cdf study report - european space agencyemits.esa.int/emits-doc/1-5200-rd19-les_total.pdfles cdf...

CCDDFF SSttuuddyy RReeppoorrtt

LLEESS

AArrcchhiitteeccttuurree SSttuuddyy ffoorr

SSuussttaaiinnaabbllee LLuunnaarr EExxpplloorraattiioonn

CDF-33(A) December 2004

This page intentionally left blank

LES CDF Study Report

Report: CDF-33(A) December 2004

page 1 of 317

s

CDF STUDY REPORT

ARCHITECTURE STUDY FOR

SUSTAINABLE LUNAR EXPLORATION



page 2 of 317

s

FRONT COVER Lunar Excursion Vehicle and Habitation Module: two vehicles composing the lunar exploration architecture



page 3 of 317

s

STUDY TEAM

This Assessment Study was performed in the ESTEC Concurrent Design Facility (CDF) by the following interdisciplinary team: TEAM LEADER A. Santovincenzo, TEC-SYE SYSTEMS U. Thomas, HME-NXO

D. Escorial, TEC-SYE T. Buechner, TEC-SYE

TRAJECTORIES A.Martinez, TEC-ECM

MISSION ANALYSIS M. Khan, OPS-GA M. Croon, OPS-GA

THERMAL J. Barbe, TEC-MCT

CONFIGURATION D. De Wilde, TEC-MCS LIFE SUPPORT L. Ordonez, TEC-MCT S. Bayon, TEC-SYE A. Rodriguez, TEC-MCT

STRUCTURE R. Westenberg, TEC-MCS HUMAN PHYSIOLOGY P. Jost, HME-GA COST T. Bieler, TEC-SYC RADIATION H. Evans, TEC-EES

RISK P. Villar, TEC-QQD COMMUNICATIONS J. Perello, TEC-ETT

MECHANISMS S. Durrant, TEC-MMM GNC B. Udrea, TEC-ECN

PROPULSION N. Boman, TEC-MPC H. Webber, TEC-MPC N. Kutufa, TEC-MPE

DHS F. Tortosa, TEC-EDD

POWER S. Zimmermann, TEC-EPS PROGRAMMATICS/AIV M. Braghin, TEC-TCC

GROUND SEGMENT AND OPERATIONS

P. Granseuer, HME-EOI D. Patterson, OPS-OSA D. Salt, OPS-OF

Under the responsibility of: S. Hovland, HME-MRH, Study Manager Visiting guests or consultants: L. Bessone, HME-AT A. Jain, HME-MRE S. Kerr, HME-HCA R. Fisackerly, DG-X The editing and compilation of this report has been provided by: M. Kruk-Strzelecki, TEC-SYE D.M. van Eck, TEC-SYE



page 4 of 317

s

This study is based on the ESA CDF Integration Design Model (IDM), which is

copyright © 2004 by ESA. All rights reserved.

Further information and/or additional copies of the report can be requested from: Scott Hovland ESA/ESTEC/HME-MRH Postbus 299 2200 AG Noordwijk The Netherlands Tel: +31-(0)71-5654023 Fax: +31-(0)71-5654437 [email protected] For further information on the Concurrent Design Facility please contact: Massimo Bandecchi ESA/ESTEC/TOS-ACD Postbus 299 2200 AG Noordwijk The Netherlands Tel: +31-(0)71-5653701 Fax: +31-(0)71-5656024 [email protected]



page 5 of 317

sTABLE OF CONTENTS

1 INTRODUCTION................................................................................................................11

1.1 Background........................................................................................................................11 1.2 Scope .................................................................................................................................11 1.3 Document structure............................................................................................................11

2 EXECUTIVE SUMMARY .................................................................................................13 2.1 Mission summary ..............................................................................................................13

3 LUNAR EXPLORATION SCENARIO ............................................................................19 3.1 Martian technology demonstration....................................................................................19 3.2 Sustainable exploration .....................................................................................................20 3.3 Lunar Base.........................................................................................................................21

4 MISSION ARCHITECTURES ..........................................................................................23 4.1 General requirements and constraints................................................................................23

4.1.1 Safety requirements .....................................................................................................23 4.1.2 Human factor requirements .........................................................................................23 4.1.3 Planetary protection.....................................................................................................28 4.1.4 Science requirements...................................................................................................28 4.1.5 Constraints ...................................................................................................................28

4.2 Architecture trade-offs.......................................................................................................28 4.2.1 General architecture issues ..........................................................................................28 4.2.2 Trade-offs ....................................................................................................................33

4.3 Architecture characteristics ...............................................................................................43 4.4 Vehicles .............................................................................................................................43

4.4.1 Overview of mission elements ....................................................................................43 4.4.2 LEV design considerations..........................................................................................44

4.5 Crew Transfer Vehicle.......................................................................................................45 4.6 Cargo vehicle.....................................................................................................................48 4.7 Preliminary propulsion system layout ...............................................................................50

4.7.1 Propulsion system adaptation ......................................................................................51 4.8 Launch and assembly strategy...........................................................................................52

4.8.1 Assembly sequence .....................................................................................................52 4.9 Travel to the lunar surface .................................................................................................57 4.10 Overall architecture performance ......................................................................................58

4.10.1 Architecture performance ............................................................................................58 4.10.2 Heavy Lift Launch Vehicle (HLLV) ...........................................................................59 4.10.3 Small reusable launcher...............................................................................................59

4.11 Abort options .....................................................................................................................60 4.12 Communication architecture .............................................................................................62

4.12.1 Requirements and design drivers.................................................................................62 4.12.2 Relay satellites trade-off..............................................................................................62 4.12.3 Communications during LEO phase ...........................................................................63 4.12.4 Baseline communication architecture..........................................................................64 4.12.5 Frequency plan ............................................................................................................66

5 MISSION ANALYSIS .........................................................................................................67 5.1 Requirements and design drivers.......................................................................................67



page 6 of 317

s5.2 Trajectories and orbits .......................................................................................................67 5.3 Baseline design..................................................................................................................67

5.3.1 Lunar orbit and rotation...............................................................................................67 5.3.2 Earth-Moon-transfer ....................................................................................................68 5.3.3 Moon-Earth-transfer ....................................................................................................70 5.3.4 LLO stability and stationkeeping ................................................................................71

5.4 Assessment of node change costs......................................................................................74 5.4.1 Surface reachability .....................................................................................................75 5.4.2 Surface descent and ascent ..........................................................................................76 5.4.3 Eclipse situation in LLO..............................................................................................79

5.5 Options ..............................................................................................................................79 5.5.1 Hub location in L1-orbit ..............................................................................................79 5.5.2 “Hops” on the lunar surface.........................................................................................82

6 LUNAR EXCURSION VEHICLE (LEV) .........................................................................83 6.1 LEV system .......................................................................................................................83

6.1.1 Requirements and design drivers.................................................................................83 6.2 LEV – descent trajectories.................................................................................................87

6.2.1 Apollo descent trajectory design .................................................................................87 6.2.2 LEV descent trajectory design.....................................................................................92 6.2.3 LEV descent trajectory baseline design.......................................................................92

6.3 LEV – landing site analysis ...............................................................................................97 6.3.1 Lunar topography ........................................................................................................97 6.3.2 Lunar craters ..............................................................................................................100

6.4 LEV – landing system .....................................................................................................102 6.4.1 Requirements and design drivers...............................................................................102 6.4.2 Assumptions ..............................................................................................................102 6.4.3 Lander stability modelling.........................................................................................103 6.4.4 Lander stability parametric analysis ..........................................................................103 6.4.5 Leg design .................................................................................................................106

6.5 LEV – configuration........................................................................................................107 6.5.1 Requirements and design drivers...............................................................................107 6.5.2 Assumptions ..............................................................................................................108 6.5.3 Baseline design..........................................................................................................109

6.6 LEV – life support ...........................................................................................................111 6.6.1 Requirements and design drivers...............................................................................111 6.6.2 Assumptions and trade-offs .......................................................................................111 6.6.3 Baseline design..........................................................................................................116 6.6.4 List of equipment.......................................................................................................117 6.6.5 Budgets ......................................................................................................................118

6.7 LEV – propulsion ............................................................................................................119 6.7.1 Requirements and design drivers...............................................................................121 6.7.2 Assumptions and trade-offs .......................................................................................121 6.7.3 Baseline design..........................................................................................................122 6.7.4 List of equipment.......................................................................................................125

6.8 LEV – GNC.....................................................................................................................126 6.8.1 Introduction ...............................................................................................................126



page 7 of 317

s6.8.2 Requirements and design drivers...............................................................................126 6.8.3 Assumptions and trade-offs .......................................................................................127 6.8.4 Baseline design..........................................................................................................128 6.8.5 List of equipment.......................................................................................................139

6.9 LEV – structures..............................................................................................................140 6.9.1 Requirements and design drivers...............................................................................140 6.9.2 Assumptions ..............................................................................................................141 6.9.3 Baseline design..........................................................................................................141 6.9.4 Budget........................................................................................................................144

6.10 LEV – communications ...................................................................................................146 6.10.1 Requirements and design drivers...............................................................................146 6.10.2 Assumptions and trade-offs .......................................................................................146 6.10.3 Baseline design..........................................................................................................149 6.10.4 List of equipment.......................................................................................................153 6.10.5 Options ......................................................................................................................154

6.11 LEV – data handling........................................................................................................154 6.11.1 Requirements and design drivers...............................................................................154 6.11.2 Main issues and proposed building blocks ................................................................155 6.11.3 Baseline design..........................................................................................................160 6.11.4 Data budgets ..............................................................................................................163 6.11.5 List of equipment.......................................................................................................165

6.12 LEV – mechanisms..........................................................................................................165 6.12.1 Requirements and design drivers...............................................................................165 6.12.2 Baseline design..........................................................................................................167

6.13 LEV – power ...................................................................................................................171 6.13.1 Requirements and design drivers...............................................................................171 6.13.2 Assumptions and trade-offs .......................................................................................172 6.13.3 Apollo design.............................................................................................................178 6.13.4 Baseline design..........................................................................................................179 6.13.5 List of equipment.......................................................................................................189 6.13.6 Option: dust removal mechanisms for solar arrays ...................................................189

6.14 LEV – thermal .................................................................................................................190 6.14.1 Requirements and design drivers...............................................................................190 6.14.2 Assumptions and trade-offs .......................................................................................191 6.14.3 Baseline design..........................................................................................................192 6.14.4 List of equipment.......................................................................................................194

7 ORBITAL HABITATION VEHICLE (HUB) ................................................................197 7.1 Hub - systems ..................................................................................................................197

7.1.1 Systems requirements ................................................................................................197 7.1.2 System design drivers................................................................................................198 7.1.3 Baseline design..........................................................................................................199 7.1.4 Mass budgets .............................................................................................................200

7.2 Hub - configuration .........................................................................................................200 7.2.1 Requirements and design drivers...............................................................................200 7.2.2 Assumptions ..............................................................................................................201 7.2.3 Baseline design..........................................................................................................201



page 8 of 317

s7.2.4 Hub configuration including two LEVs ....................................................................204

7.3 Hub - life support.............................................................................................................205 7.3.1 Requirements and design drivers...............................................................................205 7.3.2 Assumptions and trade-offs .......................................................................................206 7.3.3 Baseline design..........................................................................................................207 7.3.4 List of equipment.......................................................................................................208

7.4 Hub - propulsion..............................................................................................................210 7.4.1 Habitation module integrated propulsion system ......................................................211 7.4.2 Translunar injection propulsion system.....................................................................214 7.4.3 Lunar orbit insertion propulsion system....................................................................223 7.4.4 Trans-Earth injection propulsion module ..................................................................228 7.4.5 Conclusions ...............................................................................................................232

7.5 Hub - AOCS ....................................................................................................................232 7.5.1 Introduction ...............................................................................................................232 7.5.2 Requirements and design drivers...............................................................................233 7.5.3 Assumptions and trade-offs .......................................................................................233 7.5.4 Baseline design..........................................................................................................233 7.5.5 Equipment list............................................................................................................234

7.6 Hub - structures ...............................................................................................................234 7.6.1 Requirements and design drivers...............................................................................234 7.6.2 Assumptions ..............................................................................................................235 7.6.3 Baseline design..........................................................................................................235 7.6.4 Radiation shielding....................................................................................................236 7.6.5 Budget........................................................................................................................237

7.7 Hub - communications ....................................................................................................238 7.7.1 Requirements and design drivers...............................................................................238 7.7.2 Assumptions and trade-offs .......................................................................................238 7.7.3 Baseline design..........................................................................................................239 7.7.4 List of equipment.......................................................................................................240

7.8 Hub - data handling .........................................................................................................240 7.8.1 Requirements and design drivers...............................................................................240 7.8.2 Main issues and proposed building blocks ................................................................240 7.8.3 Baseline design..........................................................................................................240 7.8.4 List of equipment.......................................................................................................242

7.9 Hub - mechanisms ...........................................................................................................242 7.9.1 Requirements and design drivers...............................................................................242 7.9.2 Assumptions and trade-offs .......................................................................................243 7.9.3 Baseline design..........................................................................................................245 7.9.4 Budgets ......................................................................................................................250

7.10 Hub - power .....................................................................................................................251 7.10.1 Requirements .............................................................................................................251 7.10.2 Assumptions ..............................................................................................................251 7.10.3 Trade-offs between technologies for the Hub ...........................................................253 7.10.4 Baseline design..........................................................................................................260 7.10.5 List of equipments .....................................................................................................261 7.10.6 Conclusion.................................................................................................................262



page 9 of 317

s7.11 Hub – thermal ..................................................................................................................262

7.11.1 Requirements and design drivers...............................................................................262 7.11.2 Assumptions and trade-offs .......................................................................................263 7.11.3 Baseline design..........................................................................................................263

7.12 List of equipment.............................................................................................................265 8 GROUND SEGMENT AND FLIGHT OPERATIONS .................................................267

8.1 Requirements and design drivers.....................................................................................267 8.2 Assumptions and trade-offs .............................................................................................269

8.2.1 Autonomy ..................................................................................................................272 8.3 Baseline design................................................................................................................273

8.3.1 LEO Control Centre ..................................................................................................274 8.3.2 Earth-Lunar Control Centre.......................................................................................275 8.3.3 Lunar Surface Control Centre ...................................................................................275

9 RISK....................................................................................................................................277 9.1 Mission success definition, requirements and assumptions ............................................277

9.1.1 Mission success criteria .............................................................................................278 9.2 Preliminary results...........................................................................................................278 9.3 Recommendations ...........................................................................................................279

10 PROGRAMMATICS/AIV ................................................................................................281 10.1 Development....................................................................................................................281

10.1.1 Lunar Excursion Vehicle (LEV)................................................................................281 10.1.2 Hub ............................................................................................................................281 10.1.3 Propulsion Module (PM)...........................................................................................281 10.1.4 Cargo vehicle.............................................................................................................281

10.2 Ground facilities and centres ...........................................................................................281 10.3 Environmental testing......................................................................................................282

10.3.1 Hub ............................................................................................................................282 10.3.2 LEV ...........................................................................................................................282

10.4 Model philosophy............................................................................................................282 10.5 Qualification flights.........................................................................................................283

10.5.1 Hub ............................................................................................................................283 10.5.2 LEV ...........................................................................................................................283

10.6 Ground training ...............................................................................................................284 10.7 Options ............................................................................................................................284 10.8 Precursor missions...........................................................................................................284 10.9 Development schedule.....................................................................................................284

11 COST...................................................................................................................................287 11.1 Class of estimate..............................................................................................................287 11.2 Requirements and assumptions .......................................................................................287 11.3 Cost estimate methodology .............................................................................................288 11.4 Scope of the estimate.......................................................................................................289 11.5 Cost assessment ...............................................................................................................289

11.5.1 Project Office.............................................................................................................289 11.5.2 Assembly Integration and Verification, Ground Support Equipment .......................289 11.5.3 Hardware and software..............................................................................................290

11.6 Cost-risk analysis.............................................................................................................291


Report: CDF-33(A) December 2004 page 10 of 317

s12 CONCLUSIONS ................................................................................................................293

12.1 Technology development ................................................................................................294 12.2 Complementary infrastructure.........................................................................................294

13 REFERENCES...................................................................................................................295 14 ACRONYMS ......................................................................................................................301 APPENDIX A: LANDER STABILITY MODEL .............................................................307 APPENDIX B: MISSION ARCHITECTURE MODEL ..................................................315



s1 INTRODUCTION

1.1 Background In September 2004 the Aurora exploration programme performed another step in its assessment of a European roadmap for human exploration of the Moon. A proposal was made by the Aurora exploration programme to use the ESTEC Concurrent Design Facility (CDF) to perform a system trade-off among possible lunar mission architectures.

1.2 Scope As it would be impossible to study all possible options for human lunar exploration in a single CDF study, three scenarios have been set as study cases. Later a trade will have to be preformed to find the most strategically interesting option for Europe. The three scenarios are:

1. Mars Technology Demonstration (main goal of exploring the Moon will be to demonstrate technologies, elements and even operations that would be later used for Martian exploration)

2. Lunar Excursion Exploration (excursions to different lunar surface locations) 3. Lunar Exploration Base (fixed location)

It is envisioned that the final exploration programme to be implemented will be a combination of the above scenarios. Finding common elements between them would therefore increase the robustness of a future European exploration programme.

As even this proved too ambitious for the time available, the objectives of the study were therefore reduced to perform a first technical assessment of an Extended Lunar Exploration scenario, that is, to perform a system trade-off among possible mission architectures relevant to the first two lunar exploration scenarios and to provide an assessment of the vehicles most common to those architectures. The Lunar Exploration Base option would have to be performed in a later study. These studies form part of the total work plan in support of the preparation of a future European Space Exploration Programme.

The assessment study of this lunar exploration architecture consisted of 12 design sessions including three system-level preparation sessions. The results of this study are reported in this document

1.3 Document structure This document is structured so that the background to the study is described first, followed by an executive summary that gives an overview of the mission. This is followed by the chapters introducing the mission analysis and the mission architecture trade-offs.

The subsystem designs that were performed in the study are relevant to two separate vehicles: the Orbital Habitation Module (Hub) and the Lunar Excursion Vehicle (LEV). In turn, these two vehicles are split into their main subvehicles/components. There are separate chapters for all these systems, and each contains descriptions of all the relevant subsystems.

The document concludes with the chapters relevant to other overall disciplines such as operations, cost, risk, programmatics and simulation.



sDue to the different distribution requirements, only the cost assumptions (excluding figures) are given in this report. The cost assessment and assumptions made and performed as an integral part of the concurrent engineering used for this CDF study will be published as a separate document (CDF-33(B)).



s2 EXECUTIVE SUMMARY

2.1 Mission summary To define a set of elements that could be common to the two exploration scenarios chosen, the mission requirements were analysed. For the Mars Technology Demonstration mission, the main technologies chosen were long-term habitation and surface operations. This leads to an architecture having a habitation element in orbit surface landing capability. For the Lunar Excursion Exploration missions, the driving requirements were to be able to perform significantly more surface exploration (measured in surface EVA hours) than Apollo and to have the capability to reach landing zones in all major regions of the Moon. Combining these two mission types, the common elements become a habitable Hub located close to the Moon (LLO or L1) and a descent/ascent vehicle. The main parameters for these missions are shown in Table 2-1:

Mission objectives

• To perform sustainable lunar exploration, implying building the capability for several short-duration surface missions to any location on the Moon

• To land several times at different surface locations maximising the EVA time on the surface within the programme time

Necessary vehicles

• Crew Transfer Vehicle (CTV) • Lunar Excursion Vehicle (LEV) • Orbital infrastructure around lunar polar

orbit (Hub), maximum crew of six • Cargo Vehicle (CV) • Refuel Vehicle for Hub (RV) • Propulsion modules to provide needed ∆V

for TLI, LOI, TEI manoeuvres Orbit for Hub Polar Low Lunar Orbit, altitude 100 km Landing on the Moon

About once a year a crew of three can land on the surface, maximum duration 14 days

Mission architecture

Resupply About once a year the CV delivers supplies to the Hub

Launcher

• Future version of Ariane-5 for all unmanned launches: o Assumed performance: 27 tonnes to Low Earth Orbit (LEO) o 13.4 tonnes to Lunar Transfer Orbit (LTO)

• Russian Onega launcher assumed for manned CTV launch to LEO

Assembly and docking infrastructure

Assembly and launch characteristics

• Two orbital launcher assembly ‘lines’ are available in parallel

• Launch sequence on each line: eight launches per year (unmanned)



s

Docking strategy

• Docking of vehicle modules with propulsion modules in LEO

• Docking of vehicle modules with each other and Hub in LLO

Total programme duration [years] 10.75Total number of landings 13Maximum number of EVAs 182Time between first launch and first landing [months] 21Time between landings [months] 9Total number of launches 156Total mass to LEO [tonnes] 3515Average launches per landing 12

Architecture performance

Average mass per landing [tonnes] 270Start of programme Around the year 2020

LEO to LTO on FRT [km/s] 3.164Mid course manoeuvre [km/s] 0.05LTO to LLO [km/s] 0.952

∆V requirements for manned mission manoeuvres LTO to lunar surface [km/s] 2.4

LEO to LTO on Hohmann [km/s] 3.136Mid course manoeuvre [km/s] 0.05LTO to LLO [km/s] 0.868

∆V requirements for unmanned mission manoeuvres LTO to lunar surface [km/s] 2.4∆V requirement for cargo launch

Direct launch to LTO [km/s] 0.916

Hub element in LLO

Main functions

• Safe haven for surface and in-orbit crew • Docking port for arriving/departing vehicles • Orbital infrastructure for long-term habitation (for a crew of six

during 6 months) Total dry mass with margin [tonnes]

39.31

Consumables mass [tonnes] 3.41Masses

Propellant mass [tonnes] 11.15

Structure

• Two modules: one inflatable, one rigid (this report does not describe an inflatable module on the LEV)

• Structure mass rigid part with margin: 5.58 tonnes • Structure mass inflatable part with margin: 6.08

tonnes Docking Modules are docked together in LLO Radiation Radiation storm shelter in rigid Hub, can protect crew up to

4 days, yearly dose covered by Hub design Power requirement 8.1 kW on average

Design

Power Power generation Rigid Ga/As improved cells



sPower storage Li-Ion batteries Subsystem mass 2463 kg

Antennas Ka+ band, X-band (LGA and HGA), UHF Data rates Uplink: 17.5 Mbps Downlink: 20.2 Mbps AOCS equipment

Three-axis IMU, autonomous star tracker, coarse sun-sensor, RVS, DGPS Requirement (max) Six crew for 6 months in Hub Atmosphere in Hub 101.3 kPa, 21% O2/79% N2 (ISS)

Life support system Semi-closed physicochemical life support system Grey to potable water: 95% Yellow to grey water: 95% Black to grey water: 20% Solid organic waste to food: 0%

Recycling efficiencies

Oxygen: 95% Subsystem mass 10194 kg (including consumables)

Life support

Consumables mass 4044 kg Internal temperature range

18 - 27° (inhabited zone)

Internal relative humidity

25 – 70% Thermal

Radiator wing size 77 m2

AOCS

• Required ∆V = 210 m/s • Overall total wet mass during

operation assumed 110.5 tonnes • Two storable MON/MMH,

bipropellant propulsion systems each using 16 x S400-2 EADS thrusters (each 400 N thrust, Isp = 318s)

• Subsystem mass: 4194 kg

TLI propulsion module

• Two stages • 27 tonnes each • HM7-B, cryogenic engine • Turbopump-fed system • Own RCS system

Propulsion

LOI propulsion module

• One stage • 10 tonnes • Cryogenic propellants

LEV element



s

Main functions

• De-orbit, Descent and Landing (DDL) to the lunar surface • Ascent from lunar surface • Automatic Rendezvous and Docking to the Hub • Hosting the crew during DDL, surface permanence, Ascent and

Rendezvous and Docking

Assembly strategy

• AV and SHM are launched together to LEO • DM is launched separately to LEO • Propulsion modules are docked to both parts and boosted to

LLO • Final docking of all elements to the Hub in LLO

Number of crew 3Surface stay duration [days] 14Number of modules 3Lunar material to collect [kg] 30

Main characteristics

Payload capacity [kg] 260Uplink (Earth-LEV): 17.5 Mbps During descent: 1.6 Mbps

Downlink (LEV-Earth): 20.2 Mbps During descent: 6.3 Mbps Data rates

EVA suit – LEV: 12.6 Mbps both ways Communications with Earth: G/S or Moon relay satellite

• Double dish antenna (50 cm) • X- and Ka+-band • Three transponders Communication Communications with

relay in Earth orbit: TDRSS or Artemis

• Single dish antenna (40 cm) • Ka-band • Double transponder

Module description

This module will ascend from the lunar surface and dock with the Hub. Its design is a hybrid of the Apollo ascent vehicle and a Soyuz-like capsule

Module mass 9.71 tonnesModule ∆V 2220 m/s

Structure mass

1590 kg

Power requirement 18 kWhPower Power generation and

storage Regenerative fuel cells

Subsystem mass

99 kg

Data handling subsystem mass 50.8 kgAOCS equipment RDV and docking with Hub

Ambulatory medical pack Subsystem mass 401.5 kg

Life support

Total power 343 WApollo-like design with active cooling and sublimator sized for 1-day operation

Ascent Vehicle (AV)

Subsystems

Thermal

Subsystem mass: 272.4 kg



s• Rocketdyne RS-41 engines:

4x12 kN • Liquid bi-propellant

(MON/MMH) pressure-fed engines

Subsystem dry mass: 855 kg

Propulsion

Propellant mass: 4856 kgModule description

Self-standing cylindrical habitation volume, connected to the AV, launched from Earth, attached to the AV and left behind on the lunar surface

Module mass 6.85 kgHabitable volume 12.5 m3 per crew memberAtmospheric pressure 48 kPaO2 percentage in atmosphere 40%

Power requirement

457 kWh

Power generation and storage

Solar array and regenerative fuel cells

Power

Subsystem mass 233 kgStructure mass 1508 kgData handling subsystem mass 65.5 kgAOCS Docking capability with DM + AVLife support Open loop life support mass

Medical equipment as on ISS Subsystem mass 1707.4 kg Total power 635.6W

Thermal

• Deployable shielded radiator sized for the worst hot case

• Single-phase, active cooling for the cabin and the electronic equipment

Surface Habitation Module (SHM)

Subsystems

Subsystem mass 542.5 kgModule description

Purely propulsive module

Module mass 26.38 tonnesModule ∆V 2470 m/sLanding characteristics

Slope: on 7.5 km, slope shall not exceed 10 Maximum rock height/depression: 0.65 m Power Supplied by SHM

Descent Module (DM)

Subsystems Structure mass 1660 kg



s

GNC

Allows lunar landing with the following characteristics:

• Horizontal velocity 0.5 m/s • Vertical velocity 0 at 1.5 m • Manual takeover at

h=250 m Docking capability with SHM

Landing system

• Four legs • Footprint diameter 10 m • Maximum horizontal

velocity 0.5 m/s Passive design: MLI covering tanks and structures, high-temperature insulation for the main engine and heating power (50W) for thrusters and tanks

Thermal

Subsystem mass 88.7 kg• Four Aestus and 13 R-40B • 10:1 throttability (140 –14

kN) and gimballing • Hovering capability • Storable propellants

Subsystem dry mass 1780 kg

Propulsion

Propellant mass 21 635 kgTable 2-1: Executive summary



s3 LUNAR EXPLORATION SCENARIO

The ultimate goal of the Aurora Exploration programme is human exploration of the solar system and in particular of Mars. An intermediate step to achieve this goal is represented by exploration of the Moon. Three possible lunar exploration scenarios and associated high-level objectives have been considered as background for this study:

• Mars Technology Demonstration Scenario. Within this scenario, the main objective for exploring the Moon will be to demonstrate technologies and operations for later Martian exploration.

• Sustainable Exploration Scenario. The objective of this scenario will be to build the capability for several short surface missions to any location on the Moon.

• Base Scenario. The objective of this scenario will be the establishment of a continuously inhabited base at a selected location on the surface of the Moon.

From the overall objective of each scenario, a set of more detailed objectives and high-level requirements has been derived. This has been used to perform the screening of potential mission architectures based on their compliance to the objectives and their relative performance with respect to a few main “quality parameters” as described further in the report.

3.1 Martian technology demonstration Although a firm Martian exploration scenario and relevant architecture and technology are not yet available, there is common agreement on a few enabling technologies that require development and possibly flight demonstration before implementation. Among all the technologies that need to be demonstrated, a priority list has been set up. The technology demonstration has been split into two levels. The primary objectives are to:

• Demonstrate long-term habitation in a deep space environment (planet surface excluded) • Demonstrate end-to-end mission operations, in particular, surface operations

One consequence of the primary objectives is that a Habitation Module is required in any mission architecture for this scenario. This module shall be placed into an orbit so that at least the radiation environment is comparable to a mission to Mars and the module is sufficiently far from Earth that crew autonomy from ground needs to be implemented. In addition, safety features comparable to long-term deep space mission needs to be implemented and tested. A second consequence of the primary objectives is that any mission architecture for this scenario shall feature lunar landing. In particular, to test repetition of operations, landing shall take place at least twice (not necessarily at different sites). The assumed reference time frame for a human mission to Mars is the 2030 timeframe.



sThe secondary objectives are to:

• Demonstrate technologies for the Habitation Module. These would be specific technologies required for a Martian mission and in addition to the ones already included for the demonstration of the primary objectives. They include, for example, propulsion technology, zero-boil-off insulation, microgravity countermeasures as centrifuge, etc.

• Demonstrate assembly operations in LEO The achievement of the primary objectives is mandatory. The secondary objectives may be included depending on their impact on the mission architecture and on their suitability to a lunar mission. A set of high-level requirements has been derived to specify the above objectives. By long-term habitation a mission is meant that includes at least 360 days in deep space. This duration definition is the result of:

• Simulation of the duration of a round trip to Mars • Duration for which habitation volume requirements (which are time dependent) are

consistent with the ones of a Martian mission • Duration such that closed life support system is mandatory

Representativity of the environment is assumed to be achieved if the Habitation Module is placed in an orbit with an apogee upward of 90 000 km from Earth. This would give a radiation environment comparable to a deep space mission. The number of crew is not considered a parameter that needs to be reproduced faithfully at this stage. Therefore, even if Martian missions are often based on a six-crewmember team, for technology demonstration, a reduced number is still considered suitable. As a reference, a crew number of three are taken, of which two crewmembers perform the landing. A crew of six is desirable but this increases the complexity and cost of the architecture. The crew should perform at least one of the landings after long-term exposure to microgravity, that is, towards the end of the mission, to test surface operations in these conditions. Long-term permanence on the Moon is not required in this scenario. Therefore, to achieve a simplification of the mission architecture, only permanence in daylight is required (14 Earth days maximum).

3.2 Sustainable exploration A sustainable exploration encompasses repetitive visits to the Moon with the scope of accessing and analysing a widespread range of locations. The detailed surface tasks, linked to overall lunar exploration objectives (e.g. scientific) have not yet been defined at this stage and with them, mobility requirements, maximum surface area to explore, instruments to be carried, specific landing sites, etc. However, at this stage, a clear objective is the maximisation of the EVA hours on the surface. For surface operations, the following assumptions have been made:

• Mobility is required. The Apollo rover shall be assumed as baseline concerning capability and performance.



s• Samples are collected at different locations around the landing site and partially analysed

in situ. A mass of 30 kg samples per mission may be assumed. The following high-level requirements have been set within this exploration scenario to better specify the objectives:

• The mission architecture shall allow a minimum of 10 landings in 10 years • The first landing shall take place before 2020 • The mission architecture shall guarantee access to any latitude and longitude on the

surface and specifically to equator, poles and far side. • At any time during the mission, the crew shall be able to reach a radiation shelter within

one day from a solar proton event warning. This is assumed to be the best capability available from a space weather monitoring system in the time frame of the exploration.

• The surface operations shall take place in daylight. The architecture shall therefore be sized for 14 days maximum crew permanence on the lunar surface.

• The mission architecture shall allow continuous Earth/Moon and Moon orbit to surface communications capabilities (including sites located on the far side)

The size of the landing crew is a free parameter and part of the mission architecture optimisation. However, the Habitation Module shall be designed to host six crewmembers.

3.3 Lunar Base This scenario represents a higher step in technological complexity compared to the previous one and leads to different mission architecture selection criteria. For this scenario, a more detailed definition of objectives was not performed. As guideline, the base shall have the following characteristics:

• Provide long-term habitation on the surface of the Moon • Be permanently inhabited • Be resupplied from Earth, at least in the first phase (no autonomy and resource

production considered) • Be modular, so that a stepwise approach can be followed in its construction. In particular,

habitation modules for three crewmembers shall be considered as basic habitation units • Allow performance of advanced in-situ science • Allow exploitation of lunar resources in excess of the base needs • Allow exploration of large areas around the base with improved mobility capabilities

(pressurised rovers) This scenario has not been analysed further within this study.



s



s4 MISSION ARCHITECTURES

4.1 General requirements and constraints General requirements applicable to all architectures are:

• Safety requirements • Human factor requirements • Physiology requirements • Radiation requirements

4.1.1 Safety requirements

The overall safety requirement is a “cultural” choice and depends on the probability of loss of life the public opinion is ready to accept. A high-level decision between the “pioneer” (high-risk) approach and the “clerk” (low-risk) approach has to be taken. For human space flight (ISS, the Space Shuttle etc), the maximum number of acceptable failures leading to loss of life over total number of missions is generally about 1/200. However, at this stage, this number cannot be verified by analysis, therefore it has been taken only as reference. For the design, only the following requirements have been set:

• All the systems shall be made fail safe and fail operational. Whenever this is not possible abort scenarios shall be built-in

• Continuous communication shall be provided

4.1.2 Human factor requirements

4.1.2.1 Habitability

The habitability requirements can be expressed in terms of “required pressurised volume”, which, as shown in Figure 4-8, is a function of the mission duration for times less than 100 days. For longer missions, the required volume stays constant.

5 (Apollo)

20

60

30 100 60015

9

TransHabISTC IOV

75

5 (Apollo)

20

60

30 100 60015

9

TransHabISTC IOV

75

Figure 4-1: Required pressurised volume as a function of mission duration (historical data)



sBased on historical data, the required pressurised volume for the Hub was determined to be 75 m3/person. Of this, 25 m3 person was determined to be equipment-free space. In the case of the surface habitation module, a gravity of 1/6-g will be affecting the astronauts, therefore the requirements changes to free surface. Based on Earth applications (see Figure 4-2), the minimum surface area per crew has been assumed to be 3 m2. Assuming a “ceiling” of 2.3 m this lead to a minimum volume of about 7 m3 per crewmember. In all cases, a minimum airlock volume for two crewmembers has been taken as 4.5 m3, in accordance with present ISS specifications.

Surface per occupant

0.93 2.32 3.33

6.66 6.698 9

17.33

0

5

10

15

20

Fallo

ut s

helte

r

Mul

tiple

Occ

upan

cy

cell Te

nt

Trai

ler/M

obile

H

ome

Nav

al

Pre

vent

ive

Med

icin

e

Luna

r bas

e

Acc

omod

atio

n la

w

Acc

omod

atio

n la

w (3

pe

rson

s)

Surf

ace

(m2)

Figure 4-2: Surface available in the different Earth systems

4.1.2.2 Accelerations

The requirements for maximum g-loads vary depending on the body axes considered (see Figure 4-3 for body axes). Crew seats are oriented so that the g-loads during the critical phases (i.e. launch and landing. etc) are along the +Gx direction, which is a direction in which higher loads can be sustained.

along -Gz semi-axis 2.0 galong -/+Gy axis 2.4 galong +Gz semi-axis 4.0 galong -Gx semi-axis 4.5 galong +Gx semi-axis 6.0 g

Figure 4-3: Body axis direction and maximum g-loads allowed

Maximum allowable g-loads at Earth departure and Moon arrival can be seen on the right-hand side of Figure 4-3. Maximum g-loads at Moon departure and Earth arrival are lower to account



sfor crew deconditioning after experiencing microgravity for several months. Due to the lack of data for very long durations during microgravity, after a certain threshold time, the maximum g-loads have been assumed as constant (optimistic approach), i.e. 4-g along the +Gx axis. The requirements for sustained g-loads are not only a function of the direction but also depend on the time of exposure. As shown in Figure 4-4, long-duration g-load limits depend linearly on the time of exposure. The +Gx axis direction has the highest allowable loads and –Gz the one with the lowest.

Limits on Linear Acceleration for Unconditioned, Properly Restrained Crew

1.00

10.00

100.00

0.00 2.00 4.00 6.00 8.00 10.00 12.00

Time of exposure (min)

Acc

eler

atio

n, g -Gz

+/- Gy+Gz-Gx+Gx

Figure 4-4: Requirements for sustained g-loads (data not considering microgravity exposure)

The requirements on the impact g-loads (dynamic) depend again on the time of exposure to the impact. Figure 4-5 represents the tolerance to short duration –Gz accelerations (worst case).

Figure 4-5: Tolerance to short-duration –Gz accelerations



s4.1.2.3 Noise

The requirements for noise levels were established to be the following: • Maximum noise exposure for 8 hours: 84 dB (NASA-STD-3000) • Maximum noise exposure for impulse sound: 140 dB (NASA-STD-3000) • Maximum ambient noise level during daytime: 60 dB • Maximum ambient noise level during night time: 55 dB

4.1.2.4 Temperature and relative humidity

The temperature and relative humidity requirements were determined in compliance with Figure 4-6. Accordingly, the temperature in the habitable environments shall be between 18 and 27 °C at all times, for all the habitable volumes.

Figure 4-6: Temperature and humidity levels

4.1.2.5 Radiation

The requirements for radiation doses depend on various factors, such as the anatomical part exposed, age and gender of the astronauts. The limits taken as a reference for the different anatomical parts are shown in Table 4-1:

0.4-3.0 a46Careera varies with gender and age at initial exposure

0.523Annual0.2511.530 DaysBFO

Ocular LensSkin

Dose Equivalent, Gy-Eq, for Exposure Interval

0.4-3.0 a46Careera varies with gender and age at initial exposure

0.523Annual0.2511.530 DaysBFO

Ocular LensSkin

Dose Equivalent, Gy-Eq, for Exposure Interval

Table 4-1: Radiation dose limits



s4.1.2.6 Effect of microgravity

Physiological systems have to adapt to the lunar gravity environment if the exposure to microgravity is long (landing after TBD months in LLO). Since the overall surface time is limited to 14 days, lunar landings of a deconditioned crew may be very inefficient. Landing always at the beginning of a crew 6-month mission has been assumed as constraint in the mission architecture. In case a landing is performed after long-term exposure to microgravity, the reconditioning time will vary for different systems. The re-adaptation of these systems is shown in Figure 4-7. A tentative limit is shown to determine if some light physical work (walking on the surface of the Moon for example) could be performed without risk. This limit has been set after extrapolation from today’s knowledge about the return to a 1-g environment. Note that, with this extrapolation, the physical limitation would be present during the first 7 days after return to the Moon’s gravity environment.

Reconditioning after spaceflight (6 month duration) with normal rehabilitation techniques on Earth (1g)

0

1

2

3

4

5

6

7

8

9

1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 33 34 35

days postlanding

inde

x of

dec

ondi

tioni

ng neuro vestibularcardiofluidsred blood cell masslean body massbonemuscle

physical condition at which light physical work could be performed without risk

Figure 4-7: Assumed reconditioning time to gravity environment

The most critical systems are the cardiovascular system (fainting), the neuro-vestibular system (dizziness, disorientation after quick turns), and the possibility of blood anaemia (weakness). The muscles would be very weak during the first 3-4 days but would rapidly adapt. Given the maximum habitation time per crew (6 months assumed), no centrifuge would be required (ISS approach). However, the following countermeasures shall be provided on board the Hub:

• Exercise device • Pharmacological countermeasures for bone loss, radiation, space anaemia, fluid loss



s4.1.2.7 Psychology

Psychological requirements other than the ones implicitly taken into account in the volume definition have not been defined. No crew composition requirement has been considered.

4.1.3 Planetary protection

There are no specific planetary protection regulations for the Moon, both for forward and backward contamination. Therefore no cleaning procedure for the spacecraft needs to be implemented or specific countermeasures (special sealing) for the samples. Jettisoning or discarding of vehicles/elements in lunar orbit or on the lunar surface is not subject to specific restrictions either.

4.1.4 Science requirements

No specific scientific requirements have been defined for this mission. A provision for collection of about 30 kg of samples (including containers) has been assumed at each landing. 50 kg of science payload have been assumed on board the LEV. Hub science activity has been assumed to be limited to biological and physiological experiments.

4.1.5 Constraints

The following constraints apply to all scenarios and relevant architectures: • Reference dates: 2020 for Mars technology demo, 2020-2025 for exploration, 2030 for

Moon base • No nuclear propulsion shall be considered at this stage (not suitable for Europe in the

given timeframe) • No nuclear power sources for surface power production shall be considered • No in-situ resource utilisation shall be considered • Long-term permanence on the lunar surface required only in the base scenario • Launch capability:

o Assumed maximum in LEO: 27 tonnes into 400 km, corresponding to the claimed capability of several expendable launcher evolutions: Atlas V heavy, Delta IV heavy+, Angara 5, Ariane-5 27

o Assumed maximum in LTO: 13.4 tonnes o Two assembly lines working in parallel with a maximum launch rate of eight

launches per year

4.2 Architecture trade-offs Many architecture options are available for human missions to the Moon, for each of the three different scenarios proposed for the study. To define a reference, several trade-offs have been performed at architecture level, focusing on the exploration scenario. It has been shown that the results applicable to the exploration scenario are valid, with some variations, also to the technology scenario. The base scenario, however, would require a different approach.

4.2.1 General architecture issues

The general issues associated with a human mission to the Moon are:



s• Transportation/assembly of vehicles • Transfer windows • L1 use • Geographical accessibility of lunar surface • Accessibility windows for lunar surface • Return capability • Return to Earth capability

4.2.1.1 Transportation/assembly of vehicles

Taking into account the launcher capabilities assumed, 27 tonnes to LEO and 13.4 tonnes into LTO, in general it is more efficient, as regards transportation to the Moon, to launch directly into LTO than to launch first into LEO and later perform a TLI manoeuvre to get to Moon transfer orbit: 27 tonnes into LEO ⇒ 8.4 tonnes after orbit insertion into LLO 13.4 tonnes into LTO ⇒ 10.4 tonnes after orbit insertion into LLO Therefore, the following approach is recommended (as shown in Table 4-2). Whenever the vehicle/module cannot be split into elements of 10.4 tonnes, it will have to be sent from LEO and assembly will be required. Two possibilities exist: assembly in LEO and assembly in the final destination, in this case, LLO. Vehicle mass (tonnes)

Module splitting Transfer strategy

M<10.4 Single module Launch into LTO with PM for LOI 10.4<M<27 Split vehicle into modules <10.4 tonnes as

far as design allows Launch to LTO with PM for LOI

M>27

Split vehicle into modules <27 tonnes or <10.4 tonnes

Launch to LTO with PM for LOI for module < 10.4 tonnes Launch to LEO, assembly of PMs for TLI and LOI

Table 4-2: Module definitions



s

Assembly in LEOAssembly in LEO Assembly in LLOAssembly in LLO

Figure 4-8: Assembly options

In the case of LEO assembly, all the modules are launched into LEO, assembled together with the required PMs and sent to the Moon. When assembly in LLO is chosen, only the single modules with the relevant PMs are assembled in LEO. The vehicles are assembled from the modules when in LLO.

4.2.1.2 LEO to polar LLO transfer opportunities

There is a phasing required in between the LEO parking orbit and the desired final orbit around the Moon (where the orbital infrastructure is located), if a transfer orbit without plane changes is required (lower ∆V). Therefore, there will be “transfer” windows from the Earth towards the Moon. This issue is analysed in more detail in section 5.3.2.4.

4.2.1.3 L1 issue

There is the possibility of placing a vehicle at one of the libration points of the Earth- Moon system, in particular at the L1 point, the equilibrium point between Earth and the Moon.

L1

~ 326000 km~ 4 days (high thrust)


L1



Earth MoonEarth Moon

Figure 4-9: L1 configuration

See section 5.5.1 for more information on this option.



s4.2.1.4 Geographical accessibility of lunar surface

In the exploration scenario it is required to land at any point on the surface. This imposes some constraints on the location of the orbital infrastructure because the landing capabilities depend on the orbit selected. Two main options were analysed: low lunar orbit and L1.

L1L1

LLO polar orbit LLO inclined orbit LLO equatorial orbit L1 point Figure 4-10: Different orbits around the Moon

The accessibility from the different orbits is described in more detail in section 5.4.1, but in summary:

• From polar orbit, all latitudes are reachable and any longitude is revisited every 14 days • From inclined orbit, only latitudes below the inclination of the orbit are reachable,

because any longitude is revisited twice every 28 days. Depending on the inclination and the latitude the interval between passes varies, i.e. one pass after 8 days and the following one after 20 days

• From equatorial orbit, only equatorial latitudes are reachable, because any longitude is revisited every 1.5 hours

• From L1, all latitudes and longitudes are reachable at any time

4.2.1.5 Accessibility windows for lunar surface

The time on the surface will be determined by the revisit time of the orbit over the landing site. It would be possible to perform a return even if the orbit is not above the landing site, but it would imply a high cost in terms of ∆V to perform the plane change manoeuvre to catch up with the orbiter for the rendezvous and docking. The revisit time of the orbit over the landing site depends on the landing latitude and the inclination of the orbit, as shown in Figure 4-11, where the orbital plane is fixed and the lunar surface is moving under it at a rate of once every 28 days (Moon rotation period).



s

360Lon 0 30 180

Lat 0

-60

-90

60

90

50

Orbits:

Equatorial

60 deg

Polar

Moon surface moving under the orbit

28 days

Landing site5 days

14 days

360Lon 0 30 180

Lat 0

-60

-90

60

90

50

360Lon 0 30 180

Lat 0

-60

-90

60

90

50

Orbits:

Equatorial

60 deg

Polar

Moon surface moving under the orbit

28 days

Landing site5 days

14 days

Figure 4-11: Landing site revisit time

For an equatorial orbit, the interval between passes is an orbital period, about two hours. For an inclined orbit, there will be an opportunity every 28 days, Moon rotation period, plus one in between, depending on the latitude of the landing site and inclination of the orbit. For example, in the case shown in Figure 4-11, it will be a pass after 5 days, and the next one after 23 days, this being later repeated on this cycle. For a polar orbit, the repetition of the passes is every 14 days. On the other hand, a surface duration of 14 days has been set as requirement. This is the duration of a lunar day. Therefore it will be required to land at about sunrise. This imposes a new window for the landing opportunities. The Sun will have to be perpendicular to the orbital plane at landing. For a given longitude a polar orbit gives up to 27 passes per year over a certain longitude, but only five of them are during local morning (6:00 – 10:00). For polar landing sites it is possible to land and take off in every orbit (2 hours). For inclined orbits more opportunities are available, but the stay duration on the surface is always longer (into the lunar night) or shorter than 14 days.

4.2.1.6 Return to orbit capability

After the landing, the orbital plane keeps rotating with respect to the lunar surface. The rotation speed depends on the inclination of the orbit, 14 days for half a revolution in the case of polar. If an abort has to be carried out before 14 days, the orbital plane will not be aligned with the landing site, and a node change manoeuvre will be required, either for the ascent vehicle or for the orbital infrastructure.



s

Not possible

t = 0 t > 0

Not possible

t = 0 t > 0

Figure 4-12: Return opportunities from the lunar surface

4.2.1.7 Return to Earth capability

To be able to return to Earth at the minimum cost in terms of ∆V, the orbital plane will have to be perpendicular to the Earth-Moon vector, so that the orbital velocity can be used. This situation occurs every 14 days. If early return is required, a ∆V penalty will have to be paid.

TEI parallel to orbital velocity

TEI perpendicular to orbital velocity

TEI parallel to orbital velocity

TEI perpendicular to orbital velocity

Figure 4-13: Orbital plane orientation with respect to Earth

4.2.2 Trade-offs

Several trade-offs have been performed at architecture level to define the baseline architecture. The main trade-offs are shown in Table 4-3:



s

Table 4-3: Architecture trade-offs

The trade-off criteria have been defined as: • Fulfilment of the scenario requirements • Mission efficiency, defined as number of EVAs performed in 10 years, which mainly

depends on the number of EVAs per landing and the frequency of the landings

4.2.2.1 General assumptions for trade-offs

• Whenever possible, each mission element is sent directly into LTO, if not, it will be sent into LEO where it will be assembled with the propulsion modules required for TLI, LOI and TEI

• Unmanned vehicles are sent into minimum energy trajectories, manned vehicles are sent into Free Return Trajectories

• Cryogenic propulsion is used for TLI and LOI, storable for TEI, descent and ascent stages

• LEV mass of 40 tonnes, CTV mass of 14 tonnes, Hub mass of 50 tonnes taken as reference; the trade-off results remain valid when considering the mass of these vehicles

• Two independent assembly lines are available in parallel, each of them with a launch rate capability of eight launches per year

4.2.2.2 Vehicle assembly location

Some of the mission elements will have to be split into modules due to the limited performances of the launch vehicle. These modules can be either assembled in LEO and later transported to LLO or sent separately to LLO and be assembled there.



sAssembly in LEO leads to bigger composites, mission elements and the required propulsion modules for insertion into LTO and later into LLO. This in turn leads to longer assembly times and eventually to higher masses due to propellant boil-off and higher gravity losses. Assembling in LLO leads to smaller composites being sent to the Moon. This provides more flexibility in assembly planning because the two assembly lines can be used in parallel (splitting elements among assembly lines), saving time and mass.

Table 4-4: Assembly location trade-off

Assembly in LLO was selected. However, the required propulsion modules for TLI, LOI and TEI will have to be assembled to the relevant module in LEO.

4.2.2.3 Moon infrastructure

Two main types of missions were considered: 1. Excursion-type missions (no permanent infrastructure around the Moon in place): visit

several locations on the lunar surface in a consecutive series of missions without reusing elements. This way, every mission will be independent from the previous one and therefore, it will be possible to freely select landing site and orbit inclination. Each mission will be composed of a Crew Transfer Vehicle (CTV) and a Lunar Excursion Vehicle (LEV), sent separately to the Moon. They will rendezvous and dock in LLO.

2. Permanent orbital Hub: a permanent infrastructure is placed around the Moon and reused in all the missions. This infrastructure can be permanently inhabited or left alone in between landings. The crew will be exchanged by means of CTV missions, while the LEV will be sent to the orbital infrastructure in an automated way. In this scenario, it will be required to send cargo missions on a regular basis to resupply the Hub.

Preliminary sizing of both types of missions was carried out. The results are shown in Table 4-5 and Table 4-6. In the case of the Hub mission, a polar orbit around the Moon was selected as location of the infrastructure for comparison.

Parameter Unit CTV mission LEV missionTotal mass to LEO [kg] 88 135Vehicle dry mass (no propulsion system) [kg] 14 40Number of TLI PM [kg] 2 3TLI PM mass [kg] 27 27Number of LOI PM [kg] 1 1LOI PM mass [kg] 9.8 14Number of TEI PM [kg] 1 0TEI PM mass [kg] 10.6 0Total number of launches [#] 4 5

Table 4-5: Excursion type mission



s

Parameter Unit CTV mission LEV mission HUB mission CARGO missionTotal mass to LEO [kg] 88 135 176 13.4Vehicle dry mass (no propulsion system) [kg] 14 40 50 10.4Number of TLI PM [kg] 2 3 4 0TLI PM mass [kg] 27 27 27 0Number of LOI PM [kg] 1 1 1 1LOI PM mass [kg] 9.8 14 18 3Number of TEI PM [kg] 1 0 0 0TEI PM mass [kg] 10.6 0 0 0Total number of launches [#] 4 5 7 5

Table 4-6: Hub mission

In both cases, the requirement of being able to land at any place on the lunar surface is fulfilled. However, to fulfil the safety requirement of an abort from the surface at any time, it would be required to add an extra propulsion module to the CTV in the case of the excursion type mission to perform a node change manoeuvre that will allow catching up with the ascent vehicle. This propulsion module implies a mass to LEO of 175 tonnes. Since the launching frequency depends on the availability of a LEV and a CTV in lunar orbit (which itself depends on the assembly time and mass of the vehicles), the additional propulsion module downgrades the landing frequency to one landing per year. In the case of the Hub architecture, a second LEV docked to the Hub can act as lifeboat in this contingency situation, so no additional propulsion module is required. For the Hub option, the landing rate is constrained by the LEV availability to one every 7.5 months, but the first landing will not occur before 18 months from the first launch, as the permanent infrastructure has to be built. Table 4-7 shows the differences between the two options:

Capability included in CTV mission

Capability included in CTV missionReturn to Earth at any time

160 m/s per year (100 km orbit)Not requiredStationkeeping

7.5 months12 monthsTime in between landings

14.7 (9 between landings)13Average number of launches per landing [#]

266 (233.4 between landings)310Average mass to LEO per landing over 10 landings [tonnes]

18 months12 monthsTime for first landing

Storm shelter in the CTV and HUB, < 1 day access from surface

Storm shelter could be provided in the CTV, but not on the surface

Radiation

LAV performs the ascent and spare LEV perform the rescue

extra PM needed within the CTV (1500 m/s, total mass to LEO)

Return from surface at any time

Depending on the orbit inclination and orientation (but enough landing opportunities)

FreeSurface accessibility

HubExcursion typeParameters

Capability included in CTV mission

Capability included in CTV missionReturn to Earth at any time

160 m/s per year (100 km orbit)Not requiredStationkeeping

7.5 months12 monthsTime in between landings

14.7 (9 between landings)13Average number of launches per landing [#]

266 (233.4 between landings)310Average mass to LEO per landing over 10 landings [tonnes]

18 months12 monthsTime for first landing

Storm shelter in the CTV and HUB, < 1 day access from surface

Storm shelter could be provided in the CTV, but not on the surface

Radiation

LAV performs the ascent and spare LEV perform the rescue

extra PM needed within the CTV (1500 m/s, total mass to LEO)


Depending on the orbit inclination and orientation (but enough landing opportunities)

FreeSurface accessibility

HubExcursion typeParameters

Table 4-7: Infrastructure trade-off summary

The Hub architecture was selected due to the higher landing rate when the safety function is taken into account.



s4.2.2.4 Hub location

Once a Hub-based architecture has been selected, there is the possibility of placing it in LLO or in the L1 point of the Moon-Earth system. An orbit around L1 is appealing because of its accessibility both from Earth and the lunar surface at any time for/from any location. However, the required ∆V and transfer times are higher than in the case of LLO, and this implies a bigger LEV and therefore, lower landing rate. Table 4-8 shows a summary of the two options:

50 m/s per year (literature data)160 m/s per yearStationkeeping

3.88 (3.28)2.47 (2.22)∆V descent (ascent) [km/s]

19.57.5Time in between landings [months]

17.8 (17 between landings)14.7 (9 between landings)Average number of launches per landing [#]

418 (399 between landings)266 (233.4 between landings)Average mass to LEO per landing [tonnes]

30 18Time for first landing [months]

Descent in impact trajectory, strategy under investigation but uncertain

YesAbort capability during descent

Storm shelter in the hub, but 4 days away from surface (duration of ascent)

Storm shelter in the hub, < 1 day from surface

Radiation

FreePhasing of orbital nodes required. Rescue vehicle required


FreeAny location accessible but phasing with the node of the orbit and the sunlight required

Surface accessibility

L1LLO polar hubOptions

50 m/s per year (literature data)160 m/s per yearStationkeeping

3.88 (3.28)2.47 (2.22)∆V descent (ascent) [km/s]

19.57.5Time in between landings [months]

17.8 (17 between landings)14.7 (9 between landings)Average number of launches per landing [#]

418 (399 between landings)266 (233.4 between landings)Average mass to LEO per landing [tonnes]

30 18Time for first landing [months]

Descent in impact trajectory, strategy under investigation but uncertain

YesAbort capability during descent

Storm shelter in the hub, but 4 days away from surface (duration of ascent)

Storm shelter in the hub, < 1 day from surface

Radiation

FreePhasing of orbital nodes required. Rescue vehicle required


FreeAny location accessible but phasing with the node of the orbit and the sunlight required

Surface accessibility

L1LLO polar hubOptions

Table 4-8: Hub location trade-off summary

The main drawbacks of the L1 approach are the large mass needed for the LEV (which reduces the landing opportunities in 10 years) and the longer time to return to the stormshelter (located in the Hub) in the event of a radiation storm. This exceeds the requirement of 1 day if the overall ∆V is to be kept below a particular threshold. An LLO orbit for the location of the infrastructure was selected.

4.2.2.5 Lunar orbits

Different LLO orbits in terms of inclination can be selected for placing the Hub. These orbits differ as regards:

• Accessibility from a Free Return Trajectory • Surface accessibility (latitudes that can be reached) • Capability to land at sunrise • Capability to remain 14 days on the surface (the Hub shall pass over the landing site after

14 days to allow ascent) • ∆V for stationkeeping due to perturbations • Windows for return to Earth

Table 4-9 shows a comparison of the three lunar orbits:



s

0160 m/s per year160 m/s per yearStation-keeping ∆V

FreeOnly for equatorial siteYesCapability to remain 14 days on the surface

1 per orbit free1 per orbit but associated DV (TBD)

Every 14 days. Return at any time possible for 670 m/s

Return windows

FreePhasing of the orbit with terminator required. 1 possibilities per month per lat

Phasing of the orbit with terminator required. 2 possibilities per year per lat

Capability to land at sunrise

FreePossibleNot accessibleAccessibility from FRT

Free∆V for node change required∆V for node change requiredReturn from surface at any time

Nearly equatorialOther lat can be accessed by inclination changes

Only lat ≤ 33 deg Other lat can be accessed by inclination changes

Any latitudeSurface accessibility

EquatorialSun-synchronous (-33 deg)PolarOptions

0160 m/s per year160 m/s per yearStation-keeping ∆V

FreeOnly for equatorial siteYesCapability to remain 14 days on the surface

1 per orbit free1 per orbit but associated DV (TBD)

Every 14 days. Return at any time possible for 670 m/s

Return windows

FreePhasing of the orbit with terminator required. 1 possibilities per month per lat

Phasing of the orbit with terminator required. 2 possibilities per year per lat

Capability to land at sunrise

FreePossibleNot accessibleAccessibility from FRT

Free∆V for node change required∆V for node change requiredReturn from surface at any time

Nearly equatorialOther lat can be accessed by inclination changes

Only lat ≤ 33 deg Other lat can be accessed by inclination changes

Any latitudeSurface accessibility

EquatorialSun-synchronous (-33 deg)PolarOptions

Table 4-9: Moon orbit trade-off summary

The requirement to land at any latitude is the main driver, and therefore a polar orbit shall be selected. If this requirement were lifted, the best option would be the equatorial orbit. The drawback of the polar orbit is the added ∆V for return to Earth at any time (to be paid by the CTV) and the accessibility from free return trajectory. Several possibilities also exist for orbit altitude and maintenance strategy.

Parameter Polar Equatorial 60, 120 deg Sun syncronous, 143 deg

Time to periselenium altitude of 0 km [days]

180 Stable between 100 and 45 km

360 120

Node drift [deg/day] 0 NA + or – 0.6 0.986 ∆V for orbit maintenance [m/s/year]

160 0 60 260

Table 4-10: Orbit inclination around the Moon

An equatorial orbit would be preferable because the propellant mass required to maintain the orbit of the Hub is a considerable percentage of the total mass, but landing capability at any latitude is required. The required ∆V for orbit maintenance could also be lowered by raising the orbit altitude. A 200 x 200 km orbit would require virtually no ∆V for maintenance but the orbit will be elliptic during most of the time, increasing the complexity of the rendezvous and docking manoeuvres.

An intermediate solution for the orbit maintenance can be envisaged, circularising the orbit prior to every rendezvous and docking manoeuvre (up to 6 times per year). A trade-off was performed between a 100 km circular orbit maintained, a 200 km circular orbit not maintained, and a 200 km orbit circularised prior to every RdV and docking manoeuvre.



sOption Rendezvous

orbit ∆V for maintenance (m/s/year)

∆V for attitude control (m/s/year)

Total ∆V (m/s/year)

Propellant mass required (kg/year)

100x100 maintained Circular 160 50 210 9300**

200x200 not maintained Elliptical 0 50 (TBC)* 50 2100**

200x200 circularised prior to RdV Circular 77 50 (TBC)* 117.5 5700*** * ACS requirements for elliptical orbits not analysed, optimistic value ** Mass computed over an average Hub complex mass *** Mass computed manoeuvre per manoeuvre

Table 4-11: Hub orbit altitude and maintenance trade-off

Eventually, a circular orbit of 100 km altitude was retained due to its simplicity for operations, although the propellant requirements for orbit maintenance are higher.

4.2.2.6 Crew on the surface

The number of crew that will land on the surface of the Moon is a fundamental sizing parameter that will largely influence the mission efficiency. The larger the crew number, the larger the volume (and the mass) to be landed, which leads to a larger LEV and therefore longer assembly time and fewer landings in 10 years. However, a small number of crew implies lower activity while on the surface (number of EVAs). The mission efficiency has been defined as the total number of EVAs performed in 10 years:

NEVA= NEVA per landing x Nlanding in 10 yrs Where:

• the number of landings is proportional to the mass of the LEV • the number of EVAs is proportional to the number of crew, surface stay duration and

allowed activity of the astronauts. This last parameter has a great impact. Two options for EVA activity have been considered: a conservative approach based on current ISS regulations and a more ambitious one to maximise the return. The main difference between the two approaches is the rest time. In the conservative approach, it is mandatory to have at least one day rest between EVAs, while in the second approach, the astronauts are assumed to work continuously during the first four days on the surface and later are allowed to rest one day after every one or two EVAs (see Table 4-12).



s

Table 4-12: Crew activity (left: conservative approach; right: maximum return; R=rest E=EVA)

For all the possibilities of number of crew and stay duration, a preliminary sizing of the LEV was carried out which enabled defining the number of landings in 10 years. This information has been added to the number of EVAs per landing to compute the mission efficiency shown in Figure 4-14 and Figure 4-15.

0

10

20

30

40

50

60

70

0 1 2 3 4 5 6 7

Number of crew

Num

ber o

f EV

A's

in 1

0 ye

ars

14 days9 days6 days4 daysApolloReal Apollo

Figure 4-14: Mission efficiency for the conservative approach



s

0

20

40

60

80

100

120

0 1 2 3 4 5 6 7

Number of crew

Num

ber o

f EVA

's in

10

year

s

14 days9 days

6 days4 days

ApolloReal Apollo

Figure 4-15: Mission efficiency for the maximum outcome approach

Note the large difference in efficiency between the two approaches, especially for short duration missions. Concentrating on the maximum outcome approach, note that the increase in the surface stay time directly translates in a higher total number of EVAs. This is due to the relatively low impact that the stay of duration has on the vehicle sizing and therefore the number of landings. Regarding the number of crew, for short durations, a crew of two performs better than a crew of three, as the same number of EVAs are carried out with a lower mass. As the surface stay duration gets longer, this effect is compensated and overtaken as the larger the crew, the larger the number of EVAs performed. A larger number of crew (four) leads to higher number of EVAs, but when a crew of six is used, the LEV mass becomes huge and therefore the total number of landings is drastically reduced, although the total number of EVAs may increase. A crew of three for 14 days has been selected because it offers the best compromise in terms of mission efficiency and number of landings. Note also the good performance of the Apollo-type mission, with two crew for 4 days and of the real Apollo mission.

4.2.2.7 LEV reusability

Three options have been considered for the LEV design: 1. Fully expendable: a completely new LEV is used at every landing 2. Partially reusable LEV: the ascent stage is reused for later landings, any further LEV

includes wet descent stage, propellant of ascent stage and implies refuelling in lunar orbit 3. Fully reusable: the whole vehicle is reused, one single propulsion system for descent and

ascent The reusability will be meaningful if it allows performing more landings than in the expendable case. A preliminary system-level sizing of the LEV was carried out to perform the comparison.



s Fully expendable Fully reusable

Dry/payload mass [kg] 17452 17119 Propellant mass [kg] 34755 87194

Total LEV mass [kg] 59939 121963

Table 4-13: Comparison between a fully reusable and an expendable LEV

As shown in Table 4-13, a fully reusable vehicle is not recommended, as the propellant mass required is already larger than a new non-reusable LEV. This is due to the fact that the dry descent stage has to be brought back to orbit for refuelling. No extra landings are gained.

Fully expendable Partly reusable Launch mass [kg] 59939 56132

Required propellant [kg] 140765 131825 Total mass to LEO [kg] 218102 204250

Nr of launches required (theor.) [#] 8.08 7.56 Nr of launches rounded [#] 9.00 8.00

Saved mass in reusable case [%] 0.06

Table 4-14: Comparison between a partly reusable and an expendable LEV

In the case of a partly reusable LEV, some mass saving can be gained, and therefore one more landing can be carried out in 10 years. However, this would imply a more complex design, including refuelling in Moon orbit and a lifetime of the vehicle of 10 years with a likely increase of LEV dry mass (not accounted here). A fully expendable LEV was selected.

4.2.2.8 Hub habitability

The Hub can be inhabited or man-tended or even a fully robotic station. The main functions of the Hub are to offer a safe haven for the astronauts and an orbital infrastructure to perform the assembly of the vehicles. It also holds the second LEV used as lifeboat. In the case of a robotic station, the safe haven function is lost. A man-rated Hub will also not provide a long-term safe haven function. Long-term habitability of the Hub was therefore selected. The Hub will in principle be inhabited at least bridging the time of each landing (about 1 month every 7.5 months). As design point for the Hub, habitation for 6 months a year was chosen.

4.2.2.9 Reuse of surface infrastructure

The Surface Habitation Module of the LEV could, in principle, be reused, reducing the total mass of the LEV and therefore increasing the landing rate. However, this would imply that the same landing location will be visited or some mobility would have be added to the SHM. In addition, it should be designed to survive the lunar night, which would result in higher mass. Sufficient landing accuracy and control would also be required for the following LEVs. Surface infrastructure reuse was not considered appealing and was not further investigated in this study



s4.3 Architecture characteristics Summarising all trade-offs performed, the following baseline architecture has been selected:

• The architecture features a Hub in polar, 100-km circular LLO, which can be inhabited by a crew of maximum six astronauts up to 6 months. The architecture allows a new crew arrival every 9 months as well as a re-supply of the Hub with consumables.

• Several landings (more than 10 in 10 years) of three crewmembers at different lunar locations are possible, virtually at any latitude

• The maximum stay on the lunar surface per landing is 14 days. • The following vehicles, apart from the Hub, are part of the architecture infrastructure:

o Lunar Exploration Vehicle (LEV) – one per landing (about every nine months) o Crew Transportation Vehicle (CTV) – one about every nine months o Cargo Vehicle (CV) – one every year to resupply the Hub o Propulsion Modules (PM) – as required for each vehicle (or vehicle module), to

perform TLI, LOI, disposal burns and TEI (in case of the CTV) o A Refuel Vehicle (RV) that docks automatically to the Hub and refuels its tanks with

the propellant needed for stationkeeping once a year • The CTV is launched by Onega (future Russian man-rated launcher for Kliper) • Two assembly lines (and launch lines) are available and working in parallel (in a

cooperative programme this can be one for Ariane-5 and one for Delta IV super-heavy, for example)

• Both assembly lines are capable of providing a launch rate of eight launches per year • When a vehicle is split into modules, the final assembly of the vehicle is always

performed in LLO and not in LEO (only the necessary propulsion modules are added in LEO)

• The crew is transferred to the Moon on a free return trajectory • All unmanned vehicles are sent on minimum energy trajectories (Hohmann) • In the event of a crew in the Hub that prepares for a lunar landing, there are always two

LEVs docked to the Hub for safety reasons • A cargo vehicle with the Hub resupplies always arrives after the human excursion to the

lunar surface (there is always a spare docking port in case of crew arrival to the Hub) • Abort trajectories/capabilities are available for the manned vehicles

4.4 Vehicles 4.4.1 Overview of mission elements

Figure 4-16 shows the vehicles for the baseline architecture, the number of vehicles needed for the entire exploration programme, their masses and the number of required propulsion modules and launches.



s

Figure 4-16: Summary of vehicle characteristics

Some design considerations are mentioned in the following sections, for some of the four main elements and the system-level designs of the CTV and the CV are presented. Thee Hub and LEV were studied in more detail, as described in Chapter 6 Lunar Excursion Vehicle (LEV) and Chapter 7, Orbital Habitation Vehicle (Hub).

4.4.2 LEV design considerations

The efficiency of the architecture depends clearly on the landing rates that can be performed within an assumed exploration programme duration of 10 years. The landing depends on the total mass of the LEV and on the optimal splitting of masses for launch and delivery to LLO since the

Hub Total dry mass (including propellant for station-keeping): 54 t Hub configuration: 2 modules First module dry: 27 t Second module dry: 27 t PMs to deliver: 2*(27+27+9) t Total launch mass: 180 t Total nr of LEO launches: 8

LEV Total mass: 43 t Ascent Stage: 9.5 t Surface Habitation Volume: 7 t Descent Module: 26.5 t PMs to deliver: AV+SHV= 3, DM = 3 Total nr of LEO launches: 7

CTV (Kliper) Total mass: 13t PMs to deliver: 3 Number of CTVs in 10 yrs: 13 Total nr of LEO launches: 3 (plus crew on Onega)

Cargo vehicle (ATV derived):

Total mass: 13.4t (direct launch to LTO, no PMs) Number of Cargos in 10 yrs: 12



slaunch vehicle only has a capacity of 27 tonnes. An optimal splitting is not always possible or feasible because of configuration constraints. Table 4-15 shows the mass and splitting options that can be considered for the LEV:

DM mass (tonne)

Available mass for SHM and AV (tonne)

LEV total mass (tonne)

PMs for DM (LOI plusTLI,

tonne)DM launches

AV and SHM assembled in LEO

number of launches

Total number of launches

12.0 8.65 20.65 4.5+27 2 1 314.0 10.00 24.00 5+2x15 3 2 519.5 14.00 33.50 7.5+2x20 3 2 527.0 19.50 46.50 9.4+2x27 4 3 7

Table 4-15: LEV mass and splitting options

Depending on the mass of the DM, the mass of the payload for the descent (which consists of SHM and AV) can be calculated. Note that the masses increase in steps, due to the increasing number of launches required for the specific configuration. The last option with a DM of 27 tonnes was chosen, requiring a total of seven launches because this is the option that gives the highest payload mass to the surface

4.5 Crew Transfer Vehicle The Crew Transfer Vehicle (CTV) is used to transportthe crew from Earth to Moon and back. It also has some additional space for cargo. To allow an architecture mass evaluation, the Russian Kliper vehicle (see Figure 4-17) has been assumed as baseline allowing for a few modifications to adapt it to lunar missions.

Figure 4-17: Russian Kliper vehicle

The Kliper vehicle is designed to transport crew and cargo to the International Space Station. The vehicle weighs 13 tonnes and is composed of two main parts: a 9-tonne crew cabin of conical shape that acts as a wingless glider, and, attached to the rear, a smaller oval-shaped orbital module, whose design is derived from Soyuz.



sKliper is designed to carry a crew of six as well as a half tonne of cargo to the International Space Station. The Kliper design has been assumed suitable for lunar missions with the following exceptions:

• The present Kliper propulsion system has been replaced with the relevant propulsion modules to send it to the Moon and back.

• A radiation shelter of 2.1 tonnes has been added.

Figure 4-18: Onega launch vehicle with launch escape system

Figure 4-19 shows the current Kliper design, its different modules and the mass assumptions. The Launch Escape System Propulsion System and the Cone Adapter are not part of the mass budget. The Return Vehicle (mass 8.8 tonnes) is subject to minor modifications only. Eventually, the heat shield would have to be increased due to the higher re-entry velocity. The existing propulsion module, which would not be used in the case of a lunar mission, has been removed. The gained mass (2.2 tonnes) is used for the radiation shelter and larger heat shield. The habitation module (mass 2 tonnes), in which part of the crew will stay during the transfer, has not been modified but consumables have been resized according to a lunar mission.



s

Figure 4-19: Modules of the current Kliper design

At first estimation, it can be assumed that the CTV would have a mass similar to the Kliper design (13 tonnes). Propulsion modules to perform the Trans Lunar Injection (TLI) and the Lunar Orbit Insertion (LOI) need to be added in LEO before sending the vehicle to the Moon. The relevant accommodation has not been studied. Figure 4-20 shows the interior of the vehicle. The vehicle characteristics are the following:

• Crew size: up to six • Internal available volume: 20 m3 • Cargo upload capability: 500 kg • Cargo download capability: 700 kg

Figure 4-20: Interior of the Kliper vehicle

Return vehicle: 8.8 t

Habitation Soyuz: ca. 2 t

Study assumed vehicle dry mass: 13 t

(excluding Launch Escape System)

Existing propulsion module replaced and radiation shelter

added

Return vehicle: 8.8 t

Habitation Soyuz: ca. 2 t

Study assumed vehicle dry mass: 13 t

(excluding Launch Escape System)

Existing propulsion module replaced and radiation shelter

added



s4.6 Cargo vehicle The purpose of the cargo vehicle (CV) is to resupply the Hub that is orbiting around the Moon with food, water, air, and spare parts.

Figure 4-21: Automated Transfer Vehicle (ATV)

Within this study, the CV was studied as an ATV-derived vehicle to perform the necessary resupply missions to the Moon. 6 months is the stay time for a crew of six in the Hub; this corresponds to bringing 1428 kg of consumables every mission. The current ATV is composed of three main elements. These elements are retained but the propulsion system needs to be changed significantly to perform the lunar mission.



s

ICC

EPB

EAB

ICC

EPB

EAB

Figure 4-22: ATV overall design

The three main elements of the ATV are shown in Figure 4-22, namely: 1. Integrated Cargo Carrier (ICC) 2. Electronic Avionics Bay (EAB) 3. Equipped Propulsion Bay (EPB) ⇒ to be changed

The ICC contains the free space for cargo. It has been assumed that the EAB mass would not change significantly when adapted to lunar missions. The original dry mass of the ATV is 10553 kg including the propulsion system. Some dry mass reductions of the original system are possible and in particular the following components can be removed:

• Meteoroid and debris protection system of ICC (386 kg) • Meteoroid and debris protection system of EAB (334.6 kg) • ISS refuelling hardware (395.9 kg)



s• Kurs antenna (33 kg)

In total, this means a dry mass reduction of 1150 kg.

4.7 Preliminary propulsion system layout The CV can be launched in LTO. The manoeuvres to be carried out by the propulsion system of the cargo vehicle until docking with the Hub are a midcourse manoeuvre and the lunar orbit insertion. Once the cargo has been retrieved, the vehicle can be discarded, which requires another manoeuvre and some extra ∆V (see Table 4-16).

Manoeuvre Value Units∆V Earth Moon 4082 m/s

Earth to LTO 3136 m/sMidcourse manoeuvre 50 m/sLTO to LLO 866 m/sDiscarding vehicle 30 m/s

Table 4-16: Necessary ∆Vs to be provided by cargo vehicle

The propulsion system has to be designed to perform a ∆V of 946 m/s (excluding gravity losses). Taking gravity losses into account, the engine characteristics shall be considered. Four different engine options were taken into consideration and the relevant cargo mass calculated. In the following, the evaluation process is being described:

1. Direct launch capability of the considered launcher to LTO is 13.4 tonnes. 2. Dry mass of cargo vehicle is the original ATV dry mass, reduced by the above-mentioned

components and with a first estimate of a dry mass for the new propulsion system. 3. ∆V requirements are engine specific due to the gravity losses being taken into account. 4. Minimum requirement for the available supply mass are the consumables for a crew of

six staying 6 months in orbit. 5. With the above data, the available cargo capacity can be evaluated allowing a comparison

between the engine options. The four engines taken into consideration for this study are:

1. ATV main engines (storable) 2. Aestus engine (Ariane-5 upper stage, storable) 3. Vinci engine (Ariane-5 upper stage, cryogenic) 4. HM7B engine (Ariane-5 third stage, cryogenic)



sTable 4-17 shows the calculation for the different engine options: MODIFIED PROPULSION SYSTEM OPTION 1: ATV OPTION 2: Aestus OPTION 3: Vinci OPTION 4: HM7B

assumption: f = 0.08 assumption f = 0.11 assumption f = 0.11Engine mass 111 kg 280 kg 165 kgIsp 310 s 324 s 465 s 446 sNumber of engines 4 1 1 1Thrust of one engine 490 N 28000 N 180000 N 64800 NTotal thrust 1960 N 28000 N 180000 N 64800 NDelta v required (incl. grav. losses and 30 m/s) 1186 m/s 949.4 m/s 950.6 m/s 949.1 m/sf inert 0.13 0.08 0.11 0.11Total ATV dry mass (without propulsion system) 9531.1 kg 9531.10 kg 9531.10 kg 9531.10 kgPropulsion system mass (estimated) 1022.8 kg 300.88 kg 311.54 kg 322.97 kgDry mass ATV unmod (with prop sys) 10553.9 kg 9831.98 kg 9842.64 kg 9854.07 kgPossible ATV dry mass reduction 1149.50 kg 1149.50 kg 1149.50 kgDry mass ATV mod (with propulsion system) 8682.48 kg 8693.14 kg 8704.57 kgAvailable for cargo and consumables -1481 kg 1257 kg 2186 kg 2082 kgRequired consumables (6 people, 6 months) 1428 kg 1428 kg 1428 kgAdditional cargo mass (for spares etc) -170.153 kg 759 kg 655 kgLaunch mass Ariane 5 to LTO 13400 kg 13400 kg 13400 kg 13400 kg

Propellant capacity 4327.351 kg 3460.12 kg 2520.66 kg 2613.12 kg

Table 4-17: Cargo capabilities for different engine options

Note that the original version of the ATV propulsion system is not capable of providing enough thrust for a lunar mission (1960 N available) to provide enough cargo payload. The Aestus engine does not provide enough cargo capability for the mission but both the existing HM7B engine and the Vinci engine, which is under development, provide the necessary upload of cargo and additional spares. The HM7B engine has been chosen because it provides the requested cargo capabilities and is flight qualified. In addition to this cargo payload, there are 500 kg available for cargo upload on board the CTV.

4.7.1 Propulsion system adaptation

A mass estimation was done (see Table 4-18) to check if the assumed propulsion system mass and volumes are consistent with the launcher capabilities. The propulsion system would need to be changed from storable to cryogenic propellants, so the following impacts would occur:

• Change of propellant tank pressure (from 18 bar to ca. 6 bar) • Change of propellant tank material (for propellant compatibility, change from titanium to

aluminium) • Add insulation on cryogenic tanks

The mass of the cryogenic propulsion tanks including insulation has been estimated as 4% of the propellant mass. The tank is assumed to be a common bulkhead tank containing both LH2 and LO2.



sPreliminary propulsion system design

Total propellant mass 2613.12 kgDensity LO2 1142.00 kg/m3Density LH2 71.00 kg/m3Oxidiser/fuel ratio 5.14Mass LO2 2,187.53 kgMass LH2 425.59 kgVolume LO2 1.92 m3Volume LH2 5.99 m3Total volume 7.91 m3

Tank: common bulkhead tank

Tank height 1.95 mTank radius 1.14 m

Height LO2 part 0.47 mHeight LH2 part 1.48 mProp tank mass (incl insul) 104.52 kgEngine mass 165 kgComponents and He-tank 53.45 kg

Table 4-18: Preliminary mass estimation for modified propulsion system

The above-described dimensions fit into the Equipped Propulsion Bay of the current ATV design. Therefore, the result of this preliminary system level study shows that an ATV-derived cargo vehicle with a modified propulsion system can be used to provide the necessary cargo supplies.

4.8 Launch and assembly strategy This section describes the launch and assembly sequence of the baseline architecture, the masses associated and the programme schedule.

4.8.1 Assembly sequence

4.8.1.1 Assembly of the Hub

The Hub is the first vehicle to be assembled because it serves as docking station for all other vehicles and as crew habitation module in orbit. It requires a total of eight launches that are split between the two available assembly lines. The 52-tonne Hub can be split into two equally sized modules. Three separate propulsion modules, two for the Trans Lunar Injection (TLI) and one for the Lunar Orbit Insertion (LOI) manoeuvre, are added to each of the Hub modules before transfer to the Moon. The respective propulsion modules are deployed after firing, the TLI modules before arrival on the Moon, and the LOI module is put on collision course with the Moon. The two Hub modules dock automatically once they arrive in LLO. The assembly of the modules takes 6 months in LEO.



sLaunch number

Element Element mass [tonne]

Mission mass [tonne]

Launch number



1 Hub module 1 26.6 1 Hub module 2 26.62 LOI PM Hub module 1 8.6 2 LOI PM Hub module 2 8.63 TLI PM1 Hub module 1 27 3 TLI PM1 Hub module 2 274 TLI PM2 Hub module 1 27 89.2 4 TLI PM2 Hub module 2 27 89.2

Assembly line 1 Assembly line 2

Table 4-19: Launch sequence for the Hub modules

Figure 4-23 and Figure 4-25 show the modules to be launched into LEO and their assembly sequence:

Figure 4-23: Hub elements launched to LEO Figure 4-24: Hub modules in LEO

Figure 4-25 shows the sequence of events:

Figure 4-25: Hub module transfer to Moon and docking in LLO

Total number of launches

Time from first launch [months]

8 6

x 3

x 3



s4.8.1.2 Assembly of the LEV

As soon as the first eight launches for the Hub modules are completed, both assembly lines are used to assemble modules of the first two LEVs with their relevant PMs . The Ascent Vehicle and the Surface Habitation Module can be launched together. The Descent Module weighs 27 tonnes and requires a separate launch. The same strategy as for the Hub is followed. The LEV modules are shown in Figure 4-26:

Figure 4-26: Elements to be assembled for the LEV

The assembly of the Descent Module with its three propulsion modules requires one launch more than the assembly of the other LEV module, as shown in Table 4-20.

Launch number



Launch number



5 DM1 26.3 5 AV+SHM1 + LOI PM 276 LOI PM DM1 8.6 6 TLI PM1AV+SHM1 207 TLI PM1 DM1 27 7 TLI PM2 AV+SHM2 20 678 TLI PM2 DM1 27 88.9


Table 4-20: Launch sequence for the LEV elements

One assembly line will launch the necessary modules for the Descent Module, the other line the ones for the Ascent Stage and Surface Habitation Module. Once both assemblies in LEO are finalised, the Ascent Stage is sent onto a lunar transfer trajectory.

AV+SHM

TLI-PM-2

TLI-PM-1

LOI-PM

DM

TLI-PM-2

TLI-PM-1

LOI-PM



s

Figure 4-27: LEV elements in LEO

In LLO the Hub waits for the first LEV to dock (Figure 4-28). The Ascent Stage coupled with the Surface Habitation Module is docked first to the Hub. Shortly after, the Descent Stage arrives in LLO and docks directly to the SHM.

Figure 4-28: LEV modules dock to Hub in LLO

One year after the fifth launch, the Hub and the first LEV are docked in orbit and a total of 15 launches is required to achieve this configuration.

4.8.1.3 Continuation of sequence: assembly of LEV and CTV, launch of crew and cargo

After the Hub and the first LEV are docked in orbit, the main infrastructure in lunar orbit is in place. The sequence of launches and assemblies following from this point is always repeated.

x 2

x 3

DM

AS+SHM+



sThe sequence is LEV – CTV – Cargo. These vehicles are produced and launched on both assembly lines in parallel and allow the arrival of a crew in lunar orbit every nine months. The second LEV is assembled the same way as the first one. For the CTV, the necessary propulsion modules are launched to LEO, to be docked with the Kliper vehicle, which carries the crew to orbit and will only be launched at the end of the sequence. The second LEV shall be docked at the lunar Hub before the crew starts its travel to the Moon. The crew is launched with the Russian Onega launcher, and this launch is independent from the assembly sequence. After the second LEV and the crew have arrived at the Hub, a Cargo Vehicle is launched into LTO with supplies, which are retrieved by the crew on board the Hub.

Launch number



Launch number



9 AV+SHM2 + LOI PM 2310 TLI PM1 AV+SHM2 16 8 DM2 26.311 TLI PM2 AV+SHM2 16 55 9 LOI PM DM2 8.612 LOI+TEI CTV1 20.4 10 TLI PM1 DM2 2713 TLI PM1 CTV1 27 11 TLI PM2 DM2 27 88.914 TLI PM2 CTV1 27 88.4 12 CARGO1 13.4 13.4


Table 4-21: Repeatable launch sequence for LEV, CTV and Cargo Vehicle

Figure 4-29 shows the sequence repeated about every nine months:

Figure 4-29: CTV and LEV elements in LEO

The launch sequence is explained in Table 4-21 and the arrival of the different modules in LEO is shown in Figure 4-30:

x 2

x 3

x 3

CT DMAS+SHM+LOI



s

Figure 4-30: Arrival of vehicles in LLO

The number of launches that is necessary until the arrival of the first crew and first cargo supply is 26. The first crew mission is conducted 21 months after the first launch. After the first sequence, the orbital structure is in the so-called “operational configuration” (Figure 4-31). The crew can descend to the lunar surface and the rescue LEV is available. The CV would normally dock after the surface crew has returned to the Hub to ensure that there is always a spare docking port free in case of another docking port failure.

Figure 4-31: Operational configuration before lunar landing

The operational configuration has to be established prior to every other crew arrival as well. The crew time in orbit is flexible: a maximum stay of 6 months with a crew of six has been assumed.

4.9 Travel to the lunar surface The sequence of events from crew de-docking to re-docking with the Hub is:

1. Descent of LEV to surface



s2. Surface stay (14 days) 3. Ascent to Hub 4. Docking with Hub 5. Deployment of cargo and ascent stage 6. Crew can either stay or leave lunar orbit

Figure 4-32: Orbit operations, descent and crew return

4.10 Overall architecture performance Table 4-22 summarises the activities on both assembly lines that are necessary to launch and assemble the required vehicles for the lunar exploration:.

Launch number



Launch number



1 Hub module 1 26.6 1 Hub module 2 26.62 LOI PM Hub module 1 8.6 2 LOI PM Hub module 2 8.63 TLI PM1 Hub module 1 27 3 TLI PM1 Hub module 2 274 TLI PM2 Hub module 1 27 89.2 4 TLI PM2 Hub module 2 27 89.25 DM1 26.3 5 AV+SHM1 + LOI PM 276 LOI PM DM1 8.6 6 TLI PM1AV+SHM1 207 TLI PM1 DM1 27 7 TLI PM2 AV+SHM2 20 678 TLI PM2 DM1 27 88.99 AV+SHM2 + LOI PM 23

10 TLI PM1 AV+SHM2 16 8 DM2 26.311 TLI PM2 AV+SHM2 16 55 9 LOI PM DM2 8.612 LOI+TEI CTV1 20.4 10 TLI PM1 DM2 2713 TLI PM1 CTV1 27 11 TLI PM2 DM2 27 88.914 TLI PM2 CTV1 27 88.4 12 CARGO1 13.4 13.4


Table 4-22: Summary of launch sequence on both assembly lines

4.10.1 Architecture performance

The architecture as it has been proposed fulfils the requirement of more than ten landings in 10 years of exploration programme. The impact of the launch rate assumption on the architecture performance is significant. A launch rate of five launches per year decreases the efficiency by



shalf. The first two years of the programme are spent to prepare the operational configuration in LLO; no crewed mission is possible during that time.

Total programme duration 10.75 [years]Total number of landings 13 [#]Maximum number of EVA 182 [#]Time between first launch and first landing 21 [months]Time between landings 9 [months]Total number of launches 156 [#]Total mass to LEO 3515 [tonne]Average launches per landing 12 [#]Average mass per landing 270 [tonne]

Table 4-23: Architecture performance

A total number of 156 launches is required for the whole programme. Of the total number of launches, 101 launches are needed to deliver propulsion modules. Expressed in mass, this means that the total mission mass of 3581 tonnes comprises 2220 tonnes of propulsion modules. The performance is summarised in Table 4-23.

4.10.2 Heavy Lift Launch Vehicle (HLLV)

An HLLV would reduce the amount of PMs allowing direct launch into LTO of at least the LEVs and saving therefore the relevant assembly time in LEO. If the CTV is also sent with the HLLV, the number of landings can be significantly increased (but the HLLV needs to be man-rated in that case). When analysing the launch scenarios, to increase effectiveness, the performance requirements for an HLLV would be:

• 55 tonnes in LTO or two modules for the LEV: 34 tonnes and 20.3 tonnes in LTO, respectively

• 32 tonnes in LTO FRT for the CTV • Two modules of 34 tonnes in LTO for the Hub

With a performance of about 35 tonnes into LTO, the cycle LEV-CTV for landing can be performed in 4.5 months (launch rate: eight per year) instead of 9, which results in a doubling of the number of landings. The HLLV fairing dimensions have to allow the accommodation of the LEV with its LOI stage (height > 15 m). The Hub would not be suitable for a single HLLV launch anyway due to its size. To summarise, an HLLV would greatly reduce the number of launches needed, reducing the assembly time in orbit.

4.10.3 Small reusable launcher

It was assessed how a small reusable launcher would perform to fulfil this mission, under the following assumptions:

• Launch capability of 5 tonnes into LEO at a rate of 24 per year (reusable launcher type) • Same programme as in the case of “Ariane-5 – 27 tonnes”



sThe total mass into LEO of about 3500 tonnes will require 716 launches and an assembly time of 29.8 years. The real number would probably be higher because of the increased mass due to the higher number of assemblies. In addition, the configuration of each vehicle with its numerous propulsion modules is very demanding. Using a small reusable launcher can only be envisioned as a supplement to the overall programme: such a launcher could allow quick access for smaller payloads.

4.11 Abort options For each phase of the mission in which astronauts are involved, the abort capabilities have been assessed together with their relevant impact on the design and the ∆V to be paid. Only the occurrence at the same time of catastrophic failures of the Hub and an emergency during a surface stay have not been considered. The abort analysis results phase by phase are shown in Table 4-24.



sWait for the right orientation of the orbit to come back to

Earth on the CTV

None 0 14+5

Perform a node change of the orbit of the Hub to get

the right orientation and go back to Earth with the CTV

Extra PM will be required for the Hub 1500 1+5

Come back to Earth in the CTV

Extra PM will be required for the CTV to match the higher ∆V

requirements for the direct return

663 5 (TBD)

Ascent to parking orbit in the AV, rescue the astronauts with the CTV, coming back

to Earth directly

Extra PM will be required for the CTV for node change and direct return to

Earth

1500 + 663 1+5

Ascent to parking orbit in the AV, rescue the

astronauts with the 2nd LEV, come back to Hub

and return to Earth in CTV

Propulsion capabilities of the LEV consistent, Extra propellant is required for CTV to come back

directly

3000 (LEV) + 663 (CTV)

2+5

Ascent to parking orbit in the AV, Hub performs a node

change manoeuvre to match AV orbit and come back

directly to Earth in the CTV

Extra PM required for the Hub, Extra PM will be required for CTV to come

back directly

1500+663 1+5

Surface crew comes back directly to Earth (either

surface or LEO) in the AV while the Hub crew comes

back in the CTV

Extra PM required for the AV plus redesign of the capsule, Extra PM

required for CTV

663 TBD

Orbital period, everybody in the Hub

Orbital period, part of the crew on the

surface

Main flight element failure;crew life

threatening hazard


threatening hazard

Deorbit, descent and

landingMain flight element

failure;crew life threatening hazard

Abort and initiate ascent trajectory

∆V for abort lower than the one for nominal ascent

0 0

Ascent to parking orbit in the AV, perform node change to

match Hub orbit

Extra PM will be required for AV 1500 1

Ascent to parking orbit in the AV, hub performs a node change to match AV orbit

Extra PM will be required for Hub 1500 1

Send second LEV to the surface for rescue, Hub

chanbges node to enable the landing and to match the

parking orbit of the LEV

Extra PM will be required for the Hub, precision landing required for the LEV

1500+713 2

Ascent to parking orbit in the AV, CTV performs node

change to rescue astronauts and come back to Hub

Extra PM will be required for CTV 3000 2

Ascent Main flight element failure;crew life

threatening hazard

Nothing None

Rendez-vous and docking


threatening hazard

Try again Design for at least 3 attempts, account for higher ∆V

1

Trans Earth injection Main flight element failure;crew life

threatening hazard

Nothing None

Transfer to Earth Main flight element failure;crew life

threatening hazard

Deep space manoeuvre for return? Utilisation of

RCS thrusters?

Account for bigger ∆V for re targeting Included in the 50 m/s?

0

Reentry Nothing None

Surface operations Main flight element failure;crew life

threatening hazard

Table 4-24: Abort options per mission phase

The consequences for the mission architecture and vehicle design are: • Oversize of TEI propulsion module of the CTV to allow return to Earth at any time • Second LEV docked to the Hub every time a landing is performed. This vehicle will be

used to rescue the astronauts in case they have to abort from the surface by going from the Hub orbit to the ascent vehicle and back

For the following phases, there is no capability to abort:

• Ascent from lunar surface • Trans-Earth Injection • Re-entry



s4.12 Communication architecture 4.12.1 Requirements and design drivers

All vehicles shall support Tracking, Telemetry and Command (TT&C) communications during all mission phases and for any attitude: Communications availability shall be 100% for crewed spacecraft. Direct link between the different S/C shall be provided. Redundant links shall be provided so a failure is not catastrophic. Two-way ranging and Doppler capabilities are required during all mission phases. The maximum range that should be supported is 400 000 km (max. distance Earth / Moon). The telecommand (TC) and telemetry (TM) data rates shall be selectable to improve the data rate depending on the distance. The architecture shall be able to support high data rates. Data rates should be optimised by providing realistic assumptions of on-board equipment and ground segment availability. During all mission phases, data consists of housekeeping, high-quality audio and video channels, and any additional data (for example, software updates or internet access).

4.12.2 Relay satellites trade-off

To comply with the requirement of continuous communications, one or more relay satellites are needed as described in the following sections.

4.12.2.1 Hub and LEV in LLO to Earth

If no relay element is used, the worst case of communications continuity is 61% availability. The maximum blackout is 46 minutes.

4.12.2.2 LEV on lunar surface

Three main landing sites have been considered: 1. Near side: continuity depends on G/S availability. 2. Far side: no direct visibility of Earth at any time. 3. South pole: at the pole, the Earth is ± 5 degrees above the horizon. The negative

elevation causes communications blackouts that last about two weeks. To ensure continuous communications, a relay-based architecture is therefore needed. As options for the location of the relay, the following have been considered:

• L1; however, it does not cover the far side. This option was discarded. • GEO satellites; discarded since far side of the Moon would not be covered. The Hub and

LEV will have a maximum blackout of 46 minutes while orbiting the far side. • L2; this will give coverage of the Moon’s far side and the poles and is therefore the

selected option

4.12.2.3 Relay S/C orbit

In L2, two kinds of orbits can be considered: Lissajous and halo orbits. The following characteristics of the two types are: Lissajous:

• Worst case Earth occultation is 3 h. Therefore 3 h is the maximum blackout. • Up to one Earth occultation per week occurs



s• Moon occultations of the Earth occur. • Spacecraft ∆V orbit maintenance cost: 50 m/s/year.

Halo: special case of Lissajous orbit:

• Moon occultations of the Earth do not occur. • Maximum range relay-LEV on the Moon is approximately 75 000 km. • Coverage is continuous (for a minimum relay elevation of 5° over the horizon) inside:

o Latitude: +/-77°. o Longitude: +/- 69° counted from 180° meridian. For the other locations, there is coverage, but with interruptions that can last several days.

No complete coverage at either far side or poles is achievable with one relay satellite. With two relay S/C in Lissajous or halo orbit with a phase difference of 180°, there is uninterrupted coverage.

4.12.2.4 Redundancy in relay S/C

The relay system shall be two-fault tolerant. For that reason, it is necessary to have two relays for Hub and LEV at the far side of the Moon. Some short (less than 40 minutes) blackouts shall occur but Hub and LEV shall have the required autonomy to overcome them. Outside the region [lat ±77° long ±69°], continuous communications require two relay spacecraft. In case of failure of one relay, long blackouts occur. There are two solutions:

1. Add a third relay so the phase spacing is 120 degrees. This solution is expensive but in case of a single failure, the coverage of the poles is guaranteed.

2. In case of relay failure, the surface mission is aborted in advance. If the failure occurs while on the lunar surface, the other relay and the Hub, that will pass every orbit over the landing point will be used so that minimum communications will be guaranteed. This option is cheaper and since surface phase operations last 14 days, the probability of having a relay total failure is assumed to be very low.

4.12.3 Communications during LEO phase

During LEO it is difficult and expensive to provide continuous coverage using G/S due to the low orbit altitude. The solution is to use a relay satellite network, such as TDRSS. Three bands are supported by the new TDRSS H, I and J satellites:

1. S-band forward 2020.0 MHz to 2123.5 MHz, return 2200 MHz to 2300 MHz. Maximum data rate forward: 300 kbps, return 1.5 Mbps with TDRSS MA antennas. 300 kbps and 6 Mbps, respectively, for SA antenna. In Artemis 1 Mbps return link is provided.

2. Ku forward 13.775 GHz ± 1 MHz and return 15.0034 GHz ± 3 MHz. Forward: 25 Mbps, return 300 Mbps with TDRSS SA antenna.

3. Ka, 25-27 GHz forward, 23GHz return. Forward 50 Mbps, return 300 Mbps with TDRSS SA antenna.



sTo comply with the above:

• Ka band using an SA antenna has been selected for manned S/C due to high data requirements.

• For non-manned S/C, a low data rate S-band is sufficient. In addition, a back-up system is provided by using G/Ss. Omni-directional coverage shall be provided and LGAs in X-band for manned S/Cs will be used to maximise transmitted/received data rates.

4.12.4 Baseline communication architecture

Figure 4-33 to Figure 4-35 summarise the communication architecture in the nominal case and in contingency. Two cases are considered:

1. Landing site is on the Moon’s near side. 2. Landing site is on the Moon’s far side or at the poles.

Figure 4-33: Architecture for standard communications at a Moon’s near side landing site



s

Figure 4-34: Architecture for standard communications at a Moon’s far side landing site

Figure 4-35: Architecture for contingency communications at a Moon’s near side landing site



s4.12.5 Frequency plan

A list of the frequencies intervals where transmission can be located is shown in Table 4-25.

Band Transmit frequency interval Receive frequency interval4.12.5.1.1 UHF 410 -420 MHz 410 -420 MHz

S-band 2020.0-2123.5 MHz 2200 -2300 MHz 802.11g band 2.4-2.4835 GHz 2.4-2.4835 GHz

X-band 8.45-8.5 GHz 7.19-7.235 GHz Ka-band 22.55 – 23.55 GHz 25.25 – 27.5 GHz

Ka+ band 37.5-38 GHz 40-40.5 GHz Table 4-25: Frequency plan



s5 MISSION ANALYSIS

The mission analysis for the Lunar Exploration Study is concerned with the transfers from Earth to Moon and back, the study of various types of lunar orbits applicable as parking orbits for a lunar Hub and the cost and feasibility of transfers to and from the lunar surface.

5.1 Requirements and design drivers The following requirements and design drivers apply for the mission analysis work:

• The transfer orbits from the Earth to the Moon and insertion into the respective target orbits shall be studied, including launch windows, transfer cost and duration, arrival conditions and abort options

• Options for placement of a Hub in lunar orbit shall be studied, including orbital conditions, stationkeeping, surface reachability and Earth return windows

• The cost of soft-landing on the surface, including gravity losses, margins and a realistic abort scenario shall be assessed. Controllability issues shall be addressed. The launch from the surface shall be regarded. In both cases, obstacle avoidance is to be taken into account.

5.2 Trajectories and orbits The following assumptions and trade-offs are made:

• Only chemical propulsion shall be regarded for orbital transfers. No trade-off with nuclear-thermal, nuclear-electric or solar electric propulsion shall be regarded in the context of this study

• Surface landings shall take place shortly after lunar dawn; a surface mission phase must be over when the landing location approaches the dusk line

• For the descent and ascent engines, a total thrust-over-initial-mass-ratio of 3 N/kg shall be assumed, with lower values regarded as options

• The Isp for the descent and landing stages is assumed as 324 s • For the lunar descent engine, throttleability shall be assumed. For the ascent stage, this is

not required • For the Hub orbit, LLOs at an altitude of 100 km and possibly alternative altitudes shall

be regarded. A trade-off with an orbit around the L1 point shall be made • The Hub orbit shall allow for frequent access to all surface locations. For non-polar

orbits, the latitude range may have to be restricted

5.3 Baseline design 5.3.1 Lunar orbit and rotation

The Earth and Moon form a binary planetary system, rotating around a common barycentre with a sidereal period of 27.3 days. The lunar orbit is eccentric, leading to significant variations around the mean distance of 384 400 km. The inclination with respect to the ecliptic is 5.1º and the node of the orbit precesses with a period of 18.6 years. Therefore, the inclination with respect to the Earth’s Equator varies between 18.3º and 28.6º (23.5º±5.1º). As is the case with most large planetary moons, the Moon has a bound rotation with a period of 27.3 days; the same face is continuously turned towards the Earth. There is a libration due to



seccentricity and varying inclination. Therefore, not only is a significant portion of the lunar surface never visible from the Earth and vice versa, but a (sunlit and hot) lunar day is almost 2 weeks long, and so is a (dark and cold) lunar night. This imposes significant constraints on the mission design.

5.3.2 Earth-Moon-transfer

The transfer to the Moon was assumed to start from a LEO, e.g., at 200 km altitude. Different classes of transfer apply – a Hohmann transfer is mass-optimal, while a free-return-transfer has some advantages in safety, which make it appropriate for manned vehicles. The transfer cost given in the following is based on the assumption that no plane change manoeuvres need to be implemented, which implies a dedicated launch or a waiting period until the launch window opens, as discussed below. Note that the cited manoeuvre sizes are impulsive unless stated otherwise. To obtain the actual manoeuvres, the gravity losses ensuing for the selected engine thrust level must be taken into account.

5.3.2.1 Hohmann transfer

A Hohmann transfer has a typical transfer duration of around 5 days. It uses an elliptical transfer orbit where the perigee grazes the initial LEO and the apogee grazes the lunar orbit at the arrival epoch. As the Moon distance constantly changes, so will the required departure and resulting arrival speeds and corresponding TLI and LOI manoeuvre sizes. The impulsive TLI size from a 200 km LEO varies between 3.127 and 3.136 km/s. The hyperbolic arrival speed at the Moon lies between 792 and 866 m/s, leading to an LOI into a 100 km LLO of between 809 and 838 m/s. The declination of the arrival asymptote with respect to the Moon’s Equator is between ±6.6°, so the resulting lunar orbit can have any inclination between 6.6° and 173.4° without incurring a penalty, simply by targeting the arrival appropriately. Thus, polar orbits (with an orbit plane perpendicular to the lunar equator) are not a problem.

5.3.2.2 Free-return transfer

A Hohmann transfer is economical and flexible, but it is not fail-safe. If the propulsion system fails after TLI, a safe return to the Earth is not guaranteed and in all likelihood will not be possible. Conversely, a free-return transfer guarantees that even in the event of a crippling failure, the spacecraft will enter the Earth’s atmosphere only around six days after TLI without requiring any large correction manoeuvres. A free-return transfer is also an ellipse, but one with a higher apogee and therefore higher energy than that of a Hohmann transfer. Therefore, the TLI is slightly larger, with an impulsive size of 3.153 to 3.164 km/s from a 200 km LEO, around 30 m/s more than the values required for the Hohmann transfer. The typical transfer duration is reduced to around 3 days. The hyperbolic arrival velocity is 1093 – 1164 m/s, leading to a considerable increase in the impulsive LOI size, which increases by around 115 m/s to 922 – 953 m/s for a 100 km target LLO. Unlike the Hohmann transfer, a FRT cannot lead to a wide range of orbital inclinations around the Moon. The transfer aims at a near-equatorial pass at a low altitude. Here, 100 km are assumed. Only these Moon fly-by conditions will allow a safe return to the Earth in case of a failure. If the LOI is performed, the spacecraft will always be in a retrograde, near-equatorial orbit. If LOI is not or cannot be performed for some reason, the near Moon pass acts like a swing-by and redirects the spacecraft back to Earth.



s The return will then also take around 3 days and lead to entry conditions consistent with those expected for the standard Hohmann return, so the Earth entry capsule will not be subject to larger thermal fluxes or structural loads during re-entry, as described in Section 5.3.3.

5.3.2.3 Disposal of the transfer stage

Typically, the stage used for TLI is jettisoned after completion of the burn. Special care must be taken, especially when using a FRT, to prevent the stage from returning to Earth. One possibility is to retarget either the payload or the stage after separation so that the stage impacts the Moon, taking care that this does not endanger surface operations.

5.3.2.4 Transfer opportunities and initial inclination

If the inclination of the LEO is greater than the inclination of the Moon’s orbit, a transfer is possible at any day. This implies that the transfer craft is launched in a dedicated launch. Then, the launch time can be matched such that the orientation of the orbit plane is appropriate for the TLI. If the inclination is always 29°, the same ground infrastructure can be used for launches at any date. In this situation, launch opportunities exist on every day. However, for free-return transfers, the launch date may be constrained by requirements for the landing location in case of a mission abort. Each launch date leads to a fixed, but always slightly varying transfer duration. In some cases, the fail-safe return might lead to a landing over land or dangerous terrain. An in-depth study should be performed to establish whether this excludes certain launch dates or whether the problem can be addressed or mitigated in some other way. For Hohmann transfers, these constraints do not apply. Matters are more complicated if rendezvous manoeuvres are to be performed in Earth and/or Moon orbit. Then, the LEO and LLO must be carefully matched, taking into account natural orbital perturbations, to obtain frequent launch windows while avoiding costly plane change manoeuvres.

Figure 5-1: Node drift in LEO as a function of altitude and inclination

Figure 5-1 shows the oblateness-induced nodal drift per sidereal month (=1 lunar orbital revolution, 27.3 days) in LEO as function of the inclination for three orbit altitudes. The case



sconsidered here is that rendezvous manoeuvres are performed both in LEO and in LLO: Individual Hub modules are transferred to the Moon separately and mated in LLO. However, to perform the TLI and LOI, it is assumed that the modules are mated with propulsion modules in LEO. The issue is to choose the LEO inclination such that transfer opportunities will allow not just a transfer to the Moon, but a rendezvous in LLO, i.e., a given orbital plane at arrival. This requires that the LEO node drift per sidereal month is an integer multiple of 180°. As can be seen, the first such instance is at around 40° of inclination (35° for a 400 km and 43° for a 200 km LEO). There, a launch opportunity exists once per sidereal month. These inclinations should be reachable from Kourou. If this is not possible, e.g., for reasons related to stage re-entry constraints, or if compatibility with a more northerly launch site such as Baikonur is sought, another region exists around 66° – 69°, where launch opportunities repeat every two sidereal months.

5.3.2.5 Supply flights at low initial inclination

As described, with an initial inclination of 29°, a launch window to the Moon exists every day. Assuming an equatorial launch site and a launch vehicle that is optimised for low-inclination (less than 29°) launches, windows for Hohmann transfers, e.g., for supply cargo flights, are limited to the periods when the Moon is near the equatorial plane at arrival. There are two such periods per month. At other times, the transfer vehicle would either have to perform a plane change manoeuvre, which is very costly, or to insert into a non-Hohmann transfer. Thus, arrival at the Moon would be advanced, or, in the majority of the cases, delayed. This hardly increases the TLI but can raise the LOI to a similar level as that required on an FRT.

5.3.3 Moon-Earth-transfer

The return transfer should normally be a Hohmann transfer for reasons of fuel economy, so only that case was taken into account for nominal scenarios. The FRT abort trajectory, which will be employed in the case of a contingency, is not a Hohmann transfer.

5.3.3.1 Nominal Hohmann transfer

For a Hohmann transfer back to the Earth, the return vehicle, assumed to be in an LLO (typical altitude 100 km), is injected into a hyperbolic lunar escape that leaves the Moon opposite to the Moon’s direction of flight. Thus, the spacecraft will be in an ellipse that grazes the Moon orbit and has a perigee altitude distinctly inside the Earth atmosphere. The typical transfer duration is around 5 days. The hyperbolic escape velocity varies between 791 and 876 m/s, depending on the Moon distance. To reach escape, a TEI from an assumed LLO of 100 km must be imparted with an impulsive size of 808 – 837 m/s at a declination in the range of ±6.6° with respect to the lunar equator. The latitude of the re-entry point varies in the range of ±29° depending on the return epoch. It cannot easily be changed, requiring either a costly manoeuvre or a possibly lengthy delay. Conversely, the longitude of the entry point can be chosen at no extra cost simply by delaying the departure time by a few hours.



s5.3.3.2 FRT abort

The FRT abort assumes that no LOI has taken place. Then, the lunar gravity during the targeted close pass will change the orbital parameters such that instead of continuing towards the apogee, the spacecraft is heading back towards the Earth. Figure 5-2 shows a case where the abort option is exercised. The red trajectory shows the Earth-Moon transfer. The spacecraft arrives “ahead” of the Moon, passes behind it at an altitude of 100 km, and is automatically injected into a return transfer to the Earth, shown in violet. No DSMs other than minor corrections are required.

Figure 5-2: Example for free-return orbit with the abort option exercised

5.3.3.3 Transfer opportunities

As mentioned above, a Hohmann transfer requires a near-equatorial departure. From a near-equatorial orbit, such opportunities exist once every orbital revolution, i.e., every two hours. For high inclinations, this is no longer true. On a polar orbit, a least-cost escape is possible only with the orbital plane parallel to the velocity vector of the Moon orbit. This is the case twice per month. At all other times, a plane change component has to be added to the escape manoeuvre to counteract the Moon’s velocity and inject the return craft onto a trajectory that enters the Earth atmosphere. The maximum total escape velocity from a polar orbit required to ensure an Earth return at any date can be as large as 1.5 km/s, as compared to 837 m/s or less in the ideal case. Gravity losses still have to be added.

5.3.4 LLO stability and stationkeeping

Lunar orbits are subject to perturbations due to: • Non-sphericity of the lunar gravitation field • Third-body gravitational attraction through Sun and Earth



sThe resulting deviations determine orbital lifetimes as well as costs for stabilisation, which were assessed by means of numerical propagation of the following circular LLO for time spans of up to ten years:

• Altitudes: 100 km and 200 km • Inclinations: polar, equatorial, 120°, sun-synchronous

A relatively simple 4x4 gravitational model was used for the sake of computational efficiency. Numerical experiments showed that models of higher degree and order lead to different short-periodic terms in the perturbation but not to changes in the orbital lifetime. At a later point this analysis should be re-done with an advanced lunar gravity model, taking into account the lunar mascons, which currently are not yet completely mapped. The numerical propagation shows that the orbital energy (semi-major axis) is quasi-constant. The orbital inclination is a critical parameter, determining variations of both nodal line and eccentricity (the inclination itself remains quasi-constant, too). The computations show quasi-linear drifts of the orbital node at different rates depending on the inclination, while the eccentricity is subject to both short-term and long-term oscillations. An orbit is considered as decayed when the eccentricity becomes too large and the pericentre hits the surface. Figure 5-3 shows the evolution of the periselenium altitude for initially circular orbits at 100 km altitude at different inclinations. While equatorial orbits appear to remain stable for long terms, sun-synchronous orbits decay within four months.

Figure 5-3: 100 km LLO periselenium altitude evolution as function of inclination

Stationkeeping of a polar 100 km orbit is mandatory; otherwise the eccentricity build-up would drive the periselenium into the ground after around half a year. Stationkeeping consists of correcting the long-term eccentricity build-up. However, as shown in Figure 5-4, short-periodic variations of up to 20 km remain. Eliminating these is unnecessary and would be prohibitively expensive.



s

Figure 5-4: Effect of stationkeeping on polar 100 km LLO

Figure 5-5 shows the periselenium altitude evolution for a polar 200 km orbit. The eccentricity is subject to a long-periodic oscillation which leads the periselenium to first decrease down to about 90 km altitude after one year and then to increase again up to the initial altitude within another year. This pattern repeats over the observed time span of 10 years. Thus, no stationkeeping seems necessary. Note, however, that this result only holds true for the simplified model of the lunar gravity field and that refined analyses may lead to different conclusions.

Figure 5-5: Ten-year propagation of a 200 km polar LLO

Still, re-circularisation of this orbit might be advantageous from time to time for the sake of preparing for rendezvous encounters. To that end, a Hohmann transfer consisting of two manoeuvres of each up to 46 m/s (impulsive) is to be performed. Gravity losses can be minimised by splitting each manoeuvre across several consecutive orbital revolutions. Table 5-1 summarises natural orbital lifetimes and stationkeeping costs for the regarded orbits.



s

LLO altitude [km] Inclination Lifetime Stationkeeping [m/s/y]100 Equatorial > 10 years 0 100 Polar 6 months 160 100 120° 1 year 60 100 Sun-sync 4 months 260 200 Polar > 10 years 0

Table 5-1: Lifetimes and annual stationkeeping costs for different LLO types

5.4 Assessment of node change costs Rendezvous encounters require that both vehicles have equal nodes. Therefore, at least one of the vehicles must turn its node prior to docking. Node changes are expensive manoeuvres, especially when performed in vicinity of the central body. Little rotations may be performed in LLO. Large changes, however, are more effectively performed at the apocentre of an intermediate HEO, provided that the savings in the node change manoeuvre compensate the additional costs of first raising the apocentre and eventually reducing it again. Figure 5-6 shows total manoeuvre costs (impulsive) for a polar 100 km orbit. Node changes of less than about 45° are best performed in LLO. Larger turns should be performed via intermediate HEO.

Figure 5-6: Node (or Inclination) change costs for a 100-km polar LLO

In the worst case, the node must be turned by 90°. This requires three manoeuvres, each about 500 m/s impulsively, to (a) insert into HEO, (b) adjust the node, and (c) re-circularise. In case of an emergency, rendezvous might be time-critical, excluding the possibility to split any of the manoeuvres across consecutive orbital revolutions. Therefore, severe gravity losses must be taken into account resulting in total manoeuvre costs definitely beyond 1500 m/s.



s5.4.1 Surface reachability

This section describes an assessment of the number of opportunities for descending from orbit to a given landing site on the lunar surface. The following circular 100 km altitude LLO were regarded:

• Equatorial • Polar • Sun-synchronous • 120° inclined

The guidelines of the study call for the following requirements:

• Landing must occur after sunrise (6 h local time) • Departure from the surface must occur before sunset (18 h local time)

The assessment takes into account the:

• Sidereal rotation of the Moon (Tsid = 27.3 d): during surface stay, the landing site on the Moon turns beneath the orbit of the Hub at a rate of about 1° per orbital revolution.

• Nodal drift of the Hub: the line of node turns at a rate depending on inclination. • Synodic rotation of the Moon (Tsyn = 29.5 d): defines the length of the lunar day and the

illumination conditions at the landing site.

5.4.1.1 Equatorial orbit

Landing sites are restricted to a narrow belt around the equator. The belt width is determined by the hovering capacity of the LM. Opportunities of landing at sunrise occur every 29.5 days. The landing site is passed once every orbital revolution (every 118 min). Therefore, descent (and ascent) may be attempted as often as every two hours. There are no expensive node or plane change manoeuvres necessary for equatorial orbits.

5.4.1.2 Polar orbit

A polar orbit offers similar flexibility regarding descent opportunities, provided that the landing site is located close to the poles. For the polar regions, there is also one pass of the Hub over the landing site every two hours. As for non-polar latitudes however, there are significantly less descent opportunities. Since a polar orbit does not experience nodal drift, the orbit is inertially fixed. Therefore, a given landing site is passed once every sidereal period of the Moon. However, because sidereal and synodic periods differ, local times between consecutive morning passes are shifted by about 1.75 hours (equivalent to about 2 days of surface stay). Out of up to 27 passes per year, only five occur during morning (6 h – 10 h), and there are only two passes shortly after sunrise, provided that the initial conditions of the orbit are chosen appropriately. There are periods of up to five months in which no descent opportunity occurs at all.

5.4.1.3 Sun-Synchronous orbit

Sun-synchronous orbits are defined by a specific inclination for which the nodal drift equals the difference between sidereal and synodic frequency of rotation. A sun-synchronous LLO at 100 km altitude is inclined by 143° (158° for 200 km). Any given landing site in the range of accessible latitudes between ± 37° is passed once every 29.5 days at the same local time. Thus, the initial conditions of the orbit can be chosen such that there is one opportunity of descent



s(ascent) every month at 6 h (18 h). As the sun-synchronous orbit is retrograde it requires less adjustments than polar or prograde orbits when approaching from a free-return transfer trajectory.

5.4.1.4 120° inclined orbit

This orbit may be regarded as a compromise between a polar and a sun-synchronous orbit in terms of surface accessibility and number of descent opportunities. Landing site latitudes may lie between ± 60° and are therefore less restricted than those for the sun-synchronous orbit. However, the local times between morning passes are shifted by about 0.8 hours, so that only up to five passes per year occur between 6 h and 10 h. These passes occur consecutively, but afterwards, there is a period of more than 200 days during which no descent opportunities occur at all. Therefore, a 120° inclined orbit does not offer any advantage when compared to a polar orbit.

5.4.1.5 Remarks

As the Moon rotates by merely 1° per orbital revolution of the Hub (corresponding to a ground track shift of 30 km or less), for any given opportunity, descent (ascent) may be attempted more than once at consecutive orbital revolutions. The number of possible attempts is determined by the hovering capacity of the descent (ascent) stage of the LEV. A surface stay of about two weeks is the optimum from a rendezvous perspective. After 14 days, the orientation of the Hub’s nodal line has turned by half a revolution and is therefore close to its configuration by the time of descent separation. No large node change manoeuvres need to be performed. In contrast, a surface stay of about one week defines the worst case, where a nodal line rotation capability of 90° is required. This will be an important factor when sizing the vehicle to deal with contingency cases (early ascent from the surface)

5.4.2 Surface descent and ascent

5.4.2.1 The descent sequence

Here, the descent phase is regarded only down to a hovering altitude, defined as 100 m, where a zero velocity is assumed. After reaching this point, the lander will slowly settle to the landing point, while also performing lateral manoeuvres is required to achieve a desired landing location. However, the final descent from the hovering altitude to the surface is not regarded here. The following descent strategy is defined:

• A de-orbit manoeuvre of 19.5 m/s at an appropriate point on the LLO inserts into an elliptical orbit with a periselenium altitude of 15 km

• When the landing craft approaches periselenium, the braking sequence is initiated. Initially, full thrust is applied until the craft has slowed down to a limiting velocity

• At the first limiting velocity, the descent engine is throttled back to less than 70% to enhance controllability during the final portion of the descent

• As a second limiting velocity is passed, the engine is further throttled back continuously down to about 33%, which should be reached when the hovering altitude is achieved at zero velocity. This throttling-back also serves to enhance controllability and to prevent a rebounding, where the spacecraft would again inadvertently climb away from the hovering point



s The trajectories are numerically integrated assuming a nominal ratio of thrust-to-initial-mass of 3 N/kg, as for the descent module of the Apollo LM. This takes into account the gravity losses.

Figure 5-7: Altitude-over-time and FPA-over-velocity during descent burn

Figure 5-7 shows the altitude over time and the flight path angle over velocity from the start of the descent burn to the hovering point. The left-hand part assesses the obstacle clearance capabilities. The evolution of the flight path angle is important for the abort scenarios. As the FPA remains shallow until the very end of the descent, when the velocity is almost zero, there is always only a small downward component in the velocity, with a sizeable residual horizontal component. This is a safety-relevant feature with considerable importance on the abort capabilities, as will be shown.

Figure 5-8: Thrust-level-over-time and thrust-acceleration-over-velocity during descent burn

In Figure 5-8, the evolution of the thrust level during the burn is shown next to the resulting acceleration. The thrust level is given relative to the maximum thrust. It remains at a value of 1 (maximum thrust) during most of the burn; then the throttling-back can be seen. The acceleration plot shows the increase in the deceleration due to the decrease in mass (as the propellant is depleted). When the engine is throttled back, the thruster acceleration diminishes to



sapproximately the level of the lunar gravitational attraction, reached at the hovering altitude and zero-velocity. The total cost of this sequence, down to the hovering point, is 1854 m/s subject to the strategy and assumptions described above, leading to a propellant percentage of 44%. This includes the gravity losses, but not yet the margins and the amount required for the final descent to the surface. The duration of the descent burn from periselenium to hovering point is 8.4 minutes. Reducing the ratio of thrust-to-initial-mass increases manoeuvre duration and the gravity losses. Initial assessment has shown that a thrust-to-initial-mass ratio of 2.5 N/kg raises the ∆-V to the hovering point by at least 2%, a ratio of 2 N/kg by at least 5%. The descent burn durations increase by around 25 and over 50%, respectively. This requires individual verification as the throttling-back sequence also changes – at the hovering point the thrust acceleration must be equal to the gravitation.

5.4.2.2 Aborting the descent

Aborting the descent sequence shall be possible at any time. The assumed strategy for an abort is as follows:

• Always jettison the descent stage. The landing vehicle will then inevitably enter free-fall, starting from whatever conditions existed when the abort decision was taken

• The ascent stage is used for the abort. It must be ignited and brought up to full power as quickly as possible

• First, the downward motion must be stopped and a safe altitude achieved. Then, horizontal acceleration is required until a safe orbit is reached

• As shown in Figure 5-7, the FPA remains shallow almost until the hovering point is reached. The consequence is that there is a large horizontal and small downward component in the velocity, which facilitates abort. Therefore, the cost of abort, assuming that the time delay between descent stage jettison and ascent stage full power is reasonably short, does not exceed that of launch from the surface.

• The worst case is that of an abort at or near the hovering point. The initial trajectory following the de-orbit burn is inherently safe because it has a periselenium altitude of around 15 km. If for some reason the decision is taken to abandon the descent, it is enough to do nothing – the altitude will increase again as the spacecraft climbs back to the aposelenium, at which point a manoeuvre of less than 20 m/s is imparted for re-circularisation. This in itself is a very important safety feature.

5.4.2.3 Ascent with obstacle avoidance

For the ascent stage, a ratio of thrust-to-initial-mass of 3 N/kg is assumed. The engine is not throttleable, full thrust is applied throughout. The initial ascent is vertical to ensure obstacle clearance. After the vertical phase, the vehicle tilts over and follows an optimal vectored thrust profile. Initially, a 10 x 100 km orbit is targeted. The cost of reaching this orbit is 1911 m/s, including gravity losses. At the aposelenium, a circularisation burn is required. This costs 21 m/s. Margins still need to be taken into account on top of the manoeuvre budget listed above.



s

Figure 5-9: Altitude-over-distance and FPA-over-altitude for ascent burn

As shown in Figure 5-9, the ascent is initially vertical, and then gradually decreases to zero as the spacecraft tilts over. This should easily allow avoidance of any obstacle surrounding the landing site.

5.4.3 Eclipse situation in LLO

The maximum eclipse duration on a 100 km LLO is 47 minutes (for an orbital period of 118 minutes) and on a 200 km LLO, 45 minutes (period 128 minutes). This refers to the maximum traverse duration of the Moon’s shadow cone. In addition, once or twice per year, the Moon passes through the Earth’s shadow, leading to the well-known phenomenon of a lunar eclipse. Naturally, this will affect everything on the lunar surface or in a close orbit. Sometimes, such eclipses are partial. Then, as seen from the Moon, part of the solar disc remains visible throughout the eclipse, so some measure of power and heat is still received. However, it also happens that the Moon passes through the core shadow. From the Moon, the Earth appears around four times larger than the Sun, so it can effectively block the sunlight. Such an eclipse may last up to 6 hours, with up to 100 minutes spent in the totality (where virtually no sunlight is received, except for a minute percentage scattered in the Earth atmosphere), the rest of the time in the penumbra cone. These eclipses are a fact of life that must be accommodated by the Hub spacecraft’s system design. These events are entirely predictable a long time in advance. Surface operations should not be planned when an eclipse is imminent.

5.5 Options 5.5.1 Hub location in L1-orbit

The chosen baseline for the assessed orbital infrastructure is a Hub in LLO. The optional alternative is a Hub in a libration orbit around the Lagrange point L1, an unstable equilibrium point located between the Earth and Moon, around 60 000 km from the Moon. It is not practical to position a spacecraft directly in the Lagrange point. The manoeuvre cost associated with holding it there would be prohibitive. However, wide orbits around an equilibrium point are



sfeasible. Numerous concrete mission plans for spacecraft located in the L1 and especially L2 point of the Sun-Earth system exist, and some actual practical experience has already been gained. There are several distinct classes of libration orbits. A Lissajous orbit is a small-amplitude libration orbit for which the motion in and out of the Moon orbit plane is approximately harmonic, but with different frequencies, leading to a Lissajous-figure as seen from the Earth or the Moon (see Figure 5-10).

Figure 5-10: Schematic diagram of example Lissajous orbit track as seen from Moon or Earth

The Halo orbit is a special case of libration orbits with larger amplitude and identical frequencies of the in- and out-of plane motion; it appears like a wide stationary ellipse in the co-rotating frame.

Figure 5-11: Schematic diagram of an example halo orbit around the Earth-Moon-L1 point

The period for one libration orbit revolution is half the orbital period of the rotating body in the regarded system, in this case, 2 weeks. In comparison, for an orbit around the Earth-Sun-Lagrange points, the period is 6 months. In addition to the track in the y-z-plane vertical to the viewing direction from Earth or Moon, there also are significant excursions along the axis of Earth and Moon. The Lagrange points L1 and L2 are quasi-stable, a spacecraft in an orbit around these points will require periodic stationkeeping manoeuvres to preclude its escaping into a geocentric, selenocentric or even heliocentric orbit, or even an impact on the Earth or the Moon. Due to the relatively small distance from the two celestial bodies in the Earth-Moon system, this impact might occur within a matter of days after leaving the L1 orbit.



s5.5.1.1 L1 orbit dynamics and stationkeeping

The orbital dynamics of libration orbits around a Lagrange point are a complex issue that was not analysed in depth in the course of the present Lunar Exploration Study. Probably a Lissajous-type orbit would be suitable for the regarded case, though this would have to be regarded in detail. Instead, simplified approximations were made to study the accessibility of surface and Hub orbit. The stationkeeping cost was taken from the literature (RD[5]). A value of 50 m/s/year is taken.

5.5.1.2 Surface accessibility from L1 orbit

The interest in libration orbits is due to the fact that any point on the lunar surface can be reached at any time from a Hub in an orbit around L1 and vice versa. A possible direct transfer trajectory from a point on a libration orbit to the surface or back is qualitatively shown in Figure 5-12. Transfers to and from the far side are equally possible.

Figure 5-12: Schematic of libration orbit-to-surface transfer

The cost of such a surface transfer (either way) is significantly higher than for LLO. The de-orbit (or insertion, for the ascent case) burn depends on the current position on the libration orbit at the time of departure or arrival and can range from several tens of m/s to more than 300 m/s. The maximum value depends on the type of libration orbit and its amplitude. The final descent burn to achieve a soft landing on the surface has an impulsive size of 2350 m/s. The launch manoeuvre’s impulsive size is the same. Gravity losses and margins need to be added, these will significantly increase the manoeuvre size. Also, in the descent case, an added amount needs to be budgeted for hovering. The duration from de-orbit to touchdown amounts to 2-5 days, again depending on the initial position on the L1 orbit. If the de-orbit burn is very small, typically the initial motion will be slow, leading to a long total transfer. Possibly these worst cases could be amended by imparting a larger de-orbit burn in these cases and then adding a mid-course manoeuvre, if required. The transfer durations also apply for the ascent case. Here too, a reduction could be achieved at the cost of a larger ascent burn and a much larger rendezvous manoeuvre. Compared to descent from LLO there is a safety issue: The direct descent as shown above places the vehicle on an impact trajectory, which is not fail-safe. If the main engine fails, the vehicle and its crew will be lost. An alternative could be to abandon the direct descent strategy and instead opt for an approach where the de-orbit manoeuvre inserts into an orbit with a periselenium distinctly above the lunar



ssurface, an intermediate burn inserts the craft into an LLO at this orbit and the descent burn, which then would be similar to the one described in Section 5.3.3, achieves the soft landing. This would leave the impulsive manoeuvre size unchanged at 2350 m/s, but it would reduce the gravity losses and thus lead to some global improvement in the descent propellant budget. One drawback would be that the total descent duration rises. More importantly, it is not clear that if imposing the constraint that the initial trajectory, if the insertion into LLO cannot be performed, should lead back to the Hub orbit, the option of reaching any point on the surface still exists. Extensive effort should be invested into this issue, should libration orbits be chosen as baseline for a lunar exploration mission.

5.5.1.3 Transfer to L1 from the Earth and back

The transfer between Earth and Hub in L1 libration orbit is best performed via a Hohmann transfer, using an elliptical transfer orbit with the apogee altitude at the L1 distance. The impulsive TLI cost for the outbound leg from a 200 km LEO is around 3115 m/s, varying slightly due to the variations in the L1 distance. Gravity losses add to this figure. Near the apogee, the velocity must be matched with that of the Hub in its libration orbit. An exact figure cannot be given, as this varies with the inclination of the transfer orbit and the current position of the Hub in its libration orbit around L1. The impulsive insertion cost is assessed to lie between 500 and 900 m/s. Thus, in the worst case, the total transfer cost would be approximately equal to that of the transfer into LLO. If restrictions in the launch window are acceptable (to arrive at the Hub when the insertion cost is smallest) a reduction of several hundreds of m/s with respect to the transfer to LLO appears feasible. The same applies for the TEI burn. In theory, the return can be initiated at any time, unlike the LLO scenario, where even in the best case a return opportunity exists only every two hours and in the worst case, every two weeks. In practice however, apart from the dependency of the TEI cost on the phase of the libration orbit, constraints are imposed by the resulting Earth arrival location.

5.5.2 “Hops” on the lunar surface

The option of using sub-orbital trajectories to move rapidly from one surface location to another was briefly regarded. A 180° hop would constitute the worst case and allow reaching any location from anywhere else. The flight duration on such a hop is less than an hour, the maximum altitude 22 km. However, the impulsive cost is 1680 m/s for the take-off 1680 m/s for the landing (plus the usual additions). Therefore, such a hop would be only insignificantly less costly than a launch into LLO with subsequent descent. Shorter hops are faster but not much cheaper, a 90° hop lasting around half an hour and costing 2 x 1650 m/s, a 45° hop one quarter of an hour at 2 x 1550 m/s.



s6 LUNAR EXCURSION VEHICLE (LEV)

6.1 LEV system 6.1.1 Requirements and design drivers

The following main functions are allocated to the LEV: 1. De-orbit, Descent and Landing (DDL) to the lunar surface 2. Ascent from lunar surface 3. Automatic rendezvous (RdV) and docking to the Hub 4. Hosting the crew during DDL, lunar surface permanence, ascent and RdV and docking

after the surface mission Because of all these functions the LEV is the most complex and critical of all the vehicles for lunar exploration. In addition, as discussed earlier, the mass of the LEV is the main parameter affecting the efficiency of the exploration programme. The design of the LEV has been based on the following set of system requirements (Table 6-1).

Table 6-1: LEV requirements

System RequirementsLanding crew 3.00 2.00 3.00Surface stay duration 14 14 daysLunar material to be brought back 30 30 kgPayload 260 260 kgNumber of landings per year 1 2Number of EVAs per day 1 0.5 1Vehicle split in three modules: Ascent Vehicle, Surface Habitation Module (extra habitable volume), Descent ModuleReusable Ascent Vehicle

Mission constraintsOperational date 2020Launcher A5 27Descent Module mass 27.00 27.00 tonLEV diameter 4.50 4.50 m

Mission RequirementsDeltaV for de-orbit 20 20 m/s

Delta-V for descent 2450 2450 m/s

Delta-V for ascent 2220 2220 m/s

No eclipse during surface operations

Safety RequirementsRescue of the crew and/or abort of mission shall be possible during all phasesSingle failure/fault/operator error tolerance for critical hazards. Two failure/fault/operator tolerance for catastrophic hazards.

Failure detection, isolation and recovery means shall be provided (automatic and manual)

Automatic detection means for at least the following hazards:* Fire* Depressurisation* Biohazards* Atmosphere degradation conditions

* Radiation

* Temperature

* Food spoilage and water contamination

Physiology Requirementsg-loads during descent and landing 4.00 4.00 g

Habitable volume per crew member 12.50 12.50 m3Radiation Organ Specific Equivalent dose Limits (BFO)

Accute event 0.15 0.15 Sv

30 days 0.25 0.25 Sv

Year 0.50 0.50 SvCareer 1.00 1.00 Sv

Surface habitat atmospheric pressure 48.00 KpaSurface habitat oxygen percentage in the atmsphere 40.00 %

UnitsIdealMinMaxAchieved?



sThe main design drivers for the LEV are:

• Mass minimisation. As described in section 4.4.2, the exploration capability is strictly related to the mass of the LEV for a given launch performance into LEO. Lower LEV masses mean shorter assembly times in LEO and therefore, more frequent landings. Increasing the number of crew and the surface stay can also increase the architecture efficiency. However, both these parameters have a detrimental effect on the LEV mass. The optimal design point is the one in which a compromise between the three parameters is found: for a given crew size and surface stay a LEV of minimum mass shall be found. Within this study, mass minimisation has been sought through pressurised volume minimisation. By reducing the habitation volume to a minimum, a significant reduction of the landed mass and, as a consequence, of the propellant mass for descent can be achieved. For permanence on the lunar surface, a challenging value of 3 m2 per crewmember has been selected

• Launcher performance. The wet mass of the Descent Module shall be < 27 tonnes. This limits the landed mass available and forces a LEV design in at least two separate modules. An alternative would be to consider staging for the Descent Module, so increasing the landed mass capability. However, this would greatly increase the vehicle complexity and the assembly time in LEO leading to a lower architecture efficiency overall.

• Capability to land anywhere on the surface. This has a large impact on the power and thermal design that shall be made able to cope with a large range of illumination and temperature conditions. The net result is a design that is over-dimensioned for most of the missions.

• Surface permanence time. If the permanence on the surface can be limited to 2-3 days, as in the Apollo case, then, the Ascent vehicle can be used as habitation module also during the surface stay. For permanence to 14 days an extra habitation volume is required. This increases the landed mass and the overall LEV mass and forces to choose some non-optimal configuration solutions (e.g. transfer of crew from Ascent vehicle to Surface Habitation Module and back shall be guaranteed). One more reason to separate the habitation module from the Ascent Vehicle is to reduce the dry mass to be lifted from the surface to the minimum needed. This latter mass is the one that “sees” a “double delta-V” (descent-ascent). An additional complexity induced by the 14-day permanence is the need to cope with the lunar environment corresponding to dawn/dusk and noon, which affects the thermal design.

• Launcher fairing. The external diameter of the LEV is limited to 4.5 m, which results in a tall configuration in the longitudinal axis. This is demanding for the GNC system during descent.

6.1.1.1 Baseline design

Taking into account the requirement discussion above, the LEV has been designed to be composed of three separate elements:

• Lunar Ascent Vehicle (LAV), a capsule-like design with a large propulsion module. Heritage comes from the Apollo Ascent vehicle and the Soyuz capsule



s• Surface Habitation Module (SHM), a self-standing habitation volume, connected to the

LAV, launched from Earth attached to the LAV and left behind on the Moon surface after the mission.

• Descent Module (DM), a purely propulsive module in charge of the descent to the lunar surface. This module is also left behind on the Moon surface after the mission.

The three modules are stacked one on top of the other according to the mission sequence. This configuration has been found to be the one satisfying all the LEV requirements and mission operations with the lowest mass and complexity. Alternative configurations involving some kind of parallel layout for the modules with the Habitation Module horizontal and two descent modules or the ascent vehicle on the side have been briefly considered and all rejected for different reasons. The maximum mass available for the LAV and the SHM comes from the Descent Module maximum mass. The design approach has been to minimise the LAV dry mass (in turn, sizing the LAV propellant mass) so to have the maximum mass possible available for the Habitation Module. Functions and hardware not strictly needed during ascent shall be allocated to the Habitation Module, which remains on the surface of the Moon. The design of the LEV is stepwise. The simplest possible design corresponds to the Apollo one in which the Ascent Vehicle acts also as Habitation Module. This is possible only for a maximum surface stay of 3 days. Once a Surface Habitation Module is introduced, the surface permanence time becomes unimportant (provided it stays within 14 days) and this module drives entirely the LEV design. No intermediate solution exists between these two approaches. The Apollo approach (LEM) represents the “minimum” LEV design and could be used within architectures suitable for the technology scenario for which surface stay duration is not an issue. The LEV design as presented in the following sections, is compatible with the accommodation of a rover of 210 kg for limited mobility around the landing site. The detailed accommodation of the rover has not been studied but is not considered an issue at this stage.

6.1.1.2 Budgets

The mass budgets of the three LEV modules are shown in Table 6-2:



s



s

Table 6-2: LEV mass budgets

6.2 LEV – descent trajectories The objective of the analysis is to define the thrust level and attitude profile required during the lunar descent. The approach followed has been to first reproduce the Apollo trajectory to validate the model and then, to apply the principles to the design of the LEV trajectory.

6.2.1 Apollo descent trajectory design

The descent trajectory is divided into three phases (see Figure 6-1): 1. Braking: the objective is to reduce the total velocity minimising the propellant

consumption. 2. Approach: the main driver is to provide the pilot with good visual monitoring of the

approach to the lunar surface. 3. Landing: this phase provides continued visual assessment of the landing site, with

possible manual control.



s

Figure 6-1: Operational phases of Apollo-powered descent

The design of the braking phase is driven mainly by two parameters: 1. Initial altitude: the lower the better in terms of propellant consumption. But it cannot be

“too” low due to safety margins. A compromise is ~ 15 km. 2. Thrust over weight (T/WEarth): Taking into account only propellant consumption

considerations, the optimal T/WEarth is 0.7 (Figure 6-2). Nevertheless, assuming that the propellant consumed during descent is approximately half of the initial mass, and assuming that for the landing phase the T/WMoon ~ 0.75, then the throttling capability of the engine should be Tmax/Tmin =16*(T/WEarth) initial . Selecting T/WEarth = 0.7 would imply the design of an engine with Tmax/Tmin ~11, which would exceed the technology capability. For this reason, the design ratio of T/WEarth = 0.42 was selected but the final mass of the Lunar Excursion Module increased from the 11 340 kg desired to the final mass of 15 328 kg. In addition, the engine, initially designed for 46 706 N, could only deliver 43 592 N once built; therefore the real T/WEarth was eventually 0.29 (Table 6-3)

Mass (kg) Thrust (N) T/WMoon T/WEarth T/Mass (N/kg) Real 15328 43592 1.75 0.29 2.84 Design 11340 46706 2.54 0.42 4.12 Optimal 11340 77845 4.23 0.70 6.86

Table 6-3: T/W ratios of Apollo lunar excursion module



s

Figure 6-2: Characteristic velocity for fuel optimum powered descents from perigee altitude

Another important driver in the design of the Apollo nominal descent trajectory was that the engine only had throttling capability from 60 to 10% of the maximum thrust. This meant that during the braking phase where the 100% thrust was being used, there was no way to accommodate dispersions in thrust or navigation errors. Therefore, the trajectory was designed with two minutes of inefficient throttling (57 ± 3%) at the end of the braking phase, to arrive with enough accuracy to the beginning of the landing phase (see Figure 6-3):

Figure 6-3: Apollo 1 time history of thrust and attitude

The design of the Landing Phase is explained in RD[12]. The initial conditions of this phase are assumed to be an altitude of 304 m above the surface, horizontal speed of 22.82 m/s and no vertical speed. The objective is to land with a velocity of 2 m/s, while the maximum vertical speed during this phase should no exceed 6.1 m/s. In addition to this, initial pitch should be kept between –90 degrees (vertical) and –60 degrees. Figure 6-4 shows different approaches to landing trajectories. To maximise the range, the best thing is to fly vertical most of the time, only nulling the horizontal velocity at the end of the



sflight, by means of an abrupt change of attitude. On the other hand, if one wants to minimise the range, then the optimal solution is to start with the biggest pitch possible (-60 deg) and with a high thrust, so that all the horizontal velocity is nulled as soon as possible. Finally, the nominal case is ruled by a progressive and smooth decrease in the pitch. Table 6-4 shows that the ∆V required for any of the three trajectories is quite similar.

Range (m) Time (m) ∆V (m/s)

Min Range 366 113 194 Nominal 945 114 189

Max Range 2073 118 202 Table 6-4: T/W ratios of Apollo lunar excursion module

Table 6-5 summarises the total ∆V for descent according to RD[11].

Phase ∆V (fps) ∆V (m/s) ~ Prop (kg) Braking 5375 1638 6728

Approach 801 244 748 Descent 600 183 526

Total 6776 2065 8001 Table 6-5: Approximate budget of Apollo ∆V during powered descent



sPitch during hovering

-95.0

-90.0

-85.0

-80.0

-75.0

-70.0

-65.0

-60.0

-55.0

-50.0

0 20 40 60 80 100 120 140time (s)

Pitc

h (d

eg)

min rangenominalmax range

[T/W moon] during hovering

0.60

0.70

0.80

0.90

1.00

1.10

1.20

1.30

0 20 40 60 80 100 120 140time (s)

T/W


Altitude vs Range (m)

0

50

100

150

200

250

300

350

0 500 1000 1500 2000 2500Range (m)

Alti

tude

(m)


Altitude vs Time (m)

0

50

100

150

200

250

300

350

0 20 40 60 80 100 120 140Time (s)

Alti

tude

(m)


Pitch during hovering

-95.0

-90.0

-85.0

-80.0

-75.0

-70.0

-65.0

-60.0

-55.0

-50.0

0 20 40 60 80 100 120 140time (s)

Pitc

h (d

eg)


[T/W moon] during hovering

0.60

0.70

0.80

0.90

1.00

1.10

1.20

1.30

0 20 40 60 80 100 120 140time (s)

T/W


Altitude vs Range (m)

0

50

100

150

200

250

300

350

0 500 1000 1500 2000 2500Range (m)

Alti

tude

(m)


Altitude vs Time (m)

0

50

100

150

200

250

300

350

0 20 40 60 80 100 120 140Time (s)

Alti

tude

(m)


Vertical Speed (m/s)

-7.00

-6.00

-5.00

-4.00

-3.00

-2.00

-1.00

0.00

1.00

0 20 40 60 80 100 120 140time (s)

Vert

ical

spe

ed (m

/s)


Horizontal Speed (m/s)

-5.00

0.00

5.00

10.00

15.00

20.00

25.00

0 20 40 60 80 100 120 140time (s)

Hor

izon

tal s

peed

(m/s

)


Altitude vs Vertical Speed

0

50

100

150

200

250

300

350

0.00 1.00 2.00 3.00 4.00 5.00 6.00 7.00Vertical Speed (m/s)

Alti

tude

(m)


Altitude vs Speed

0

50

100

150

200

250

300

350

0.0 5.0 10.0 15.0 20.0 25.0Speed (m/s)

Alti

tude

(m)


Vertical Speed (m/s)

-7.00

-6.00

-5.00

-4.00

-3.00

-2.00

-1.00

0.00

1.00

0 20 40 60 80 100 120 140time (s)

Vert

ical

spe

ed (m

/s)


Horizontal Speed (m/s)

-5.00

0.00

5.00

10.00

15.00

20.00

25.00

0 20 40 60 80 100 120 140time (s)

Hor

izon

tal s

peed

(m/s

)


Altitude vs Vertical Speed

0

50

100

150

200

250

300

350

0.00 1.00 2.00 3.00 4.00 5.00 6.00 7.00Vertical Speed (m/s)

Alti

tude

(m)


Altitude vs Speed

0

50

100

150

200

250

300

350

0.0 5.0 10.0 15.0 20.0 25.0Speed (m/s)

Alti

tude

(m)


Figure 6-4: Apollo landing phase – minimum range, nominal and maximum range trajectories



s6.2.2 LEV descent trajectory design

Compared to the Apollo trajectory, it was assumed that the descent engine would have throttle capability for the required range of thrust. In addition to this, the development of GNC systems allows to make manoeuvres with rates up to 1.5 deg/s in the last part of the descent, which will improve the overall propellant consumption. Therefore, in the LEV case, the trajectory can be split into two phases. The first one is the braking phase with 100% thrust (this maximum thrust is a parameter to be optimised) and constraints to be applied at the end (see Table 6-6):

Altitude (m) Flight Path Speed (m/s) Vertical Speed (m/s) Pitch (degrees)

In the range (100,500) In the range (20,65) 0 In the range (-80, -40)

Table 6-6: Constraints in the final conditions at the end of the braking phase of the LEV

The second phase, landing phase, should end with a pitch equal to –90 degrees (vertical landing) and flight for a range of 1 km (crater size to be avoided), to allow manoeuvrability and selection of the final landing site. The Isp of the engine is assumed to be 324 s.

6.2.3 LEV descent trajectory baseline design

Before calculating the complete trajectory, some parametric studies were performed to assess the influence of defined parameters on the two flight phases.

6.2.3.1 Braking phase

Several computations were run to verify the influence of the T/WEarth. . The final conditions were fixed and equal to the values shown in RD[12]: Final altitude = 305 m, Final Horizontal Velocity = 22.86 m/s, Final Vertical Velocity =0 m/s. The objective is to minimise the propellant consumption. The results (Table 6-7 and Figure 6-5) show that the T/W has little impact on the propellant consumption (1200 kg, which is 5.5 % of the total propellant mass). In any case, the bigger the thrust, the shorter the firing time and therefore the lower the consumption.

Initial Mass (kg) Thrust (N) T/WEarth Propellant (kg) ∆V (m/s) Time(s) Range (km) 50000 98067 0.200 21935 1834.9 711 633 50000 140000 0.286 21146 1746.8 480 432 50000 343233 0.700 20728 1701.1 192 175

Table 6-7: Influence of T/W ratio on the propellant consumption of the LEV



s

0 500 10000

5

10

15

20

Time [s]

Alti

tude

[km

]

0 500 1000-40

-20

0

20

Time [s]

Pitc

h An

gle

[deg

]

0 500 1000-150

-100

-50

0

50

Time(s)

Vert

ical

Spe

ed (

m/s

)

0 500 10000

500

1000

1500

2000

Time(s)

Hor

izon

tal S

peed

(m/s

)

T/W=0.20T/W=0.29T/W=0.70

0 1 20

5

10

15

20

Flight-Path Speed [km/s]

Alti

tude

[km

]

0 5 10

x 105

0

5

10

15

20

Horizontal Range [m]

Alti

tude

[km

]

0 500 10000

100

200

300

400

Time(s)

Thru

st (

kN)

0 500 10000

0.5

1

1.5

2

2.5x 10

4

Time(s)

Prop

ella

nt(k

g)

T/W=0.20T/W=0.29T/W=0.70

Figure 6-5: LEV braking phase for different T/W ratios

6.2.3.2 Landing phase

The initial altitude and horizontal speed are the parameters to be optimised for minimising the propellant consumption. The final velocity as well as the initial vertical velocity are constrained



sto be equal to 0 m/s, the final pitch equal to –90 degrees (vertical landing) and the range equal to 1 km. The results are shown in Figure 6-6:

0 20 40 600

0.05

0.1

0.15

0.2

Time [s]

Altit

ude

[km

]

0 20 40 60-90

-80

-70

-60

-50

Time [s]

Pitc

h An

gle

[deg

]

0 20 40 60-5

-4

-3

-2

-1

0

Time(s)

Ver

tical

Spe

ed (m

/s)

0 20 40 600

10

20

30

40

50

Time(s)

Hor

izon

tal S

peed

(m/s

)

0 0.02 0.04 0.06-0.05

0

0.05

0.1

0.15

0.2


Altit

ude

[km

]

0 500 1000-0.05

0

0.05

0.1

0.15

0.2


Altit

ude

[km

]

0 20 40 6020

30

40

50

60

Time(s)

Thru

st (k

N)

0 20 40 600.4

0.6

0.8

1

1.2

1.4

Time(s)

Thru

st to

wei

ght [

-]

Figure 6-6: LEV landing phase – 1 km range

The optimal initial conditions are initial altitude 159 m and horizontal speed of 41.45 m/s. The thrust level is also optimised, with an upper boundary of 55 kN so the maximum T/WMoon



sachieved is around 1.2, similar to the Apollo values. These conditions yield to a minimum consumption of 855 kg, equivalent to a ∆V = 94.92 m/s. The optimal parameters identified by these analyses have then been used as guideline into the definition of the nominal descent trajectory

6.2.3.3 Nominal trajectory

For the overall trajectory, the optimisable parameters are the pitch and the thrust profiles. And the objective is to minimise the overall fuel consumption. An upper limit was defined for the thrust. This upper limit was agreed to be such that the T/WMoon would be equal to 1.75, as in the Apollo mission. Fixing an upper limit for the thrust, the thrust profile during the braking phase becomes constant and equal to the maximum value. The thrust profile of the landing phase is also almost constant, so, for simplicity it was decided to make it constant, though the constant value would be optimisable. Finally, the pitch profile was modified so that the maximum rate would not be higher than 1.5 degrees/s. Table 6-8 and Figure 6-7 summarise the baseline descent trajectory:

Time (s)

Range (km)

Alt (km)

Mass (kg)

Propellant (kg)

∆V (m/s)

Pitch (degrees)

Thrust (N)

T/WMoon

Breaking 0.0 0.000 15.000 50000.0 0.0 0.0 0.70 140000 1.75 479.6 440.263 0.197 28866.2 21133.8 1745.5 -42.49 140000 2.99Landing 537.2 441.263 0.000 27903.9 962.2 107.7 -90.00 53149 1.17

Table 6-8: Influence of T/W ratio on the propellant consumption of the LEV



s

0 1 2-5

0

5

10

15

20


Alti

tude

[km

]

0 2 4 6

x 105

-5

0

5

10

15

20


Alti

tude

[km

]

0 200 400 60040

60

80

100

120

140

Time(s)

Thru

st (

kN)

0 200 400 6001

1.5

2

2.5

3

Time(s)

Thru

st to

wei

ght [

-]

0 200 400 6000

5

10

15

Time [s]

Alti

tude

[km

]

0 200 400 600

-80

-60

-40

-20

0

Time [s]

Pitc

h An

gle

[deg

]

0 200 400 600-80

-60

-40

-20

0

20

Time(s)

Verti

cal S

peed

(m/s

)

0 200 400 6000

500

1000

1500

2000

Time(s)

Hor

izon

tal S

peed

(m/s

)

Figure 6-7: Baseline trajectory



s6.3 LEV – landing site analysis To assess the terrain characteristics at landing and provide boundaries for the design of the landing system, an analysis of the lunar terrain, based on existing remote sensing data, has been performed.

6.3.1 Lunar topography

The current surface topography data source is the Clementine Instrument Data. Two data sets with different data resolutions are available:

• 1° resolution (30.334 km) • 0.25° resolution (7.5 km)

Note that due to the function used to process the data into a rectangular grid, the following data above 78°N and below 78°S are not reliable. Using the 0.25° resolution data, the surface slope between data points has been calculated in the longitude and latitude directions. These individual slopes have been averaged together and plotted on the data point in question. The slope data has then been processed to identify all areas with slopes of less than 10° inclination and overlaid upon the topography data to identify the potential landing areas and the altitude range to which the landing system needs to be designed. Due to the large amount of data, the surface has been split into four quadrants as shown in Figure 6-8:

Q1

Q3

Q2

Q4

0 180 360 -90

+90

0

Figure 6-8: Surface splitting for topography analysis

Per quadrant, the following figures show the surface slopes and landing areas with slopes less than 10 degrees which is considered achievable by the landing system



s

Max. Altitude = 4706.5 m Min. altitude = -6011.6 m

Figure 6-9: Topography data for quadrant 1

Max. altitude = 3442.6 m Min. altitude = -4643.6 m For Information (yellow markers in fig.): Apollo 11 Landing Latitude: 0.67N Longitude: 203.49 (23.49E) Apollo 15 Landing Latitude: 26.11N Longitude: 183.66 (3.66E) Apollo 17 Landing Latitude: 20.17N Longitude: 210.8 (30.8E)




s


Max. altitude = 2882.5 m Min. altitude = -5135.6 m For Information (yellow markers in fig.): Apollo 14 Landing Latitude: 3.67S Longitude: 197.46 (17.46W) Apollo 16 Landing Latitude: 8.60S Longitude: 195.31 (15.31W)


Max. altitude = 3275.9 m Min. altitude = -5634.9 m For Information (yellow markers in fig.): Apollo 12 Landing Latitude: 2.94S Longitude: 156.55 (23.49W)



page 100 of 317

sFrom Figure 6-9 to Figure 6-12, the following observations can be made:

• 7.5 km data resolution is insufficient to determine detailed surface slopes • Large areas of the surface are not accessible when considering slopes of less than 10° • Within the assessed accessible landing areas, the surface altitude is between +4.706 km

and –6.011 km The landing system design has been based on these ranges. Note that this method of graphing the results may lead to an overestimation of the non-accessible areas. The figures above should currently only be used as an indication of the accessible landing areas. To conclude, more detailed topographical information is needed to properly assess the accessible landing sites.

6.3.2 Lunar craters

No ‘Crater Model’ seems to be available that gives the probability of a certain size crater within any landing area, although data (diameter & grid reference) on individual (named) craters are available (in total 1562) from RD[15]. This is typically for craters of above 0.5 km in radius. Seven craters are listed in the reference as having diameters of 0 km, which is interpreted as having a diameter < 0.5 km. Figure 6-13 shows the distribution of the craters w.r.t. the diameter:

Total

0

5

10

15

20

25

30

0

1.8 5 10 15 20 25 30 34 39 44 49 53 58 63 68 73 78 83 88 93 98 103

108

113

118

123

128

134

141

149

158

165

177

184

190

199

219

237

272

319

Total

Count of Diameter

Diameter

Figure 6-13: Lunar crater distribution as function of their size

The largest crater listed is the ‘Hertzsprung’ crater at 591 km in diameter. There are 61 craters listed with diameters equal to or below 2 km. These are:



page 101 of 317

s 0 km Diameter Latitude Longitude 1 km Diameter Latitude Longitude 2 km Diameter Latitude Longitude[Chekov] -6.6 82 Ango 20.5 -32.3 Akis 20 -31.8[Homer] -24.3 133.6 Annegrit 29.4 -25.6 Alan -10.9 -6.1[Lorca] 24.4 10.9 Bawa -25.3 102.6 Amontons -5.3 46.8[Montesquieu] -6.2 92.3 Boris 30.6 -33.5 Artemis 25 -25.4[Sophocles] -21.5 119.8 Charles 29.9 -26.4 Christel 24.5 11[Vergil] -26.3 133 Courtney 25.1 -30.8 Collins 1.3 23.7[Voltaire] -11.9 100.3 Felix 25.1 -25.4 Curtis 14.6 56.6 Grace 14.2 35.9 Delia -10.9 -6.10.5 km Diameter Ian 25.7 -0.4 Diana 14.3 35.7Dag 18.7 5.3 Isabel 28.2 -34.1 Donna 7.2 38.3Manuel 24.5 11.3 Isis 18.9 27.5 Eckert 17.3 58.3Osama 18.6 5.2 Jerik 18.5 27.6 Freud 25.8 -52.3 Linda 30.7 -33.4 Gaston 30.9 -340.8 km Diameter Mary 18.9 27.4 Harold -10.9 -6Louise 28.5 -34.2 Mavis 29.8 -26.4 José -12.7 -1.6 Osiris 18.6 27.6 Julienne 26 3.21.5 km Diameter Robert 19 27.4 Linné 27.7 11.8Abetti 20.1 27.8 Rosa 20.3 -32.3 Monira -12.6 -1.7 Sampson 29.7 -16.5 Osman -11 -6.21.8 km Diameter -10.9 -6.2 Stella 19.9 29.8 Pupin 23.8 -11Priscilla Susan -11 -6.3 Samir 28.5 -34.3 Walter 28 -33.8 Sita 4.6 120.8 Yoshi 24.6 11 Soraya -12.9 -1.6 Vera 26.3 -43.7 Verne 24.9 -25.3

Table 6-9: Craters with size below 2 km

In general, there are two types of lunar surface crater: 1. Simple Impact Craters. Simple impact craters have bowl-shaped depressions, mostly with

smooth walls. This type of crater generally has a diameter less than 15 km. Their depth is about 20% of the diameter.

Figure 6-14: Simple impact craters

2. Complex Impact Craters. Complex impact craters have a single or multiple peaks in the middle of the crater. These craters have diameters between about 12 and 20 and 175 km, and the central uplift is usually one or a few peaks. Craters with a diameter over 175 km can have more complex, ring-shaped uplifts within the crater.



page 102 of 317

s

Figure 6-15: Complex impact craters

Additionally, there are surface impact basins: An impact basin is an impact crater that has a rim diameter greater than 300 km. There are over 40 impact basins on the Moon. These catastrophic impacts produce faulting and other crust deformations. Material ejected from impact basins is distributed over wide areas. As crater avoidance capability a size of 2 km diameter has been given as requirement for the lander GNC subsystem.

6.4 LEV – landing system 6.4.1 Requirements and design drivers

The following requirements for the landing system have been given: • The landing system shall be designed to land on a terrain with the following

characteristics: o An average slope, measured on a 7.5 km scale not to exceed 10° o An average rock height/depression depth of 0.65 m (taken from Apollo)

• The DLS shall limit the vertical landing velocity to 2.5 m/s at surface contact (soft landing). This derives to 0 m/s at max. 1.95 m above the surface for engine cut-off

• The DLS shall limit the horizontal landing velocity to 1 m/s at surface contact or Engine cut-off

• The DLS shall be designed such that craters of diameter 2 km can be avoided. • The DLS shall be designed to land on the lunar surface between – 6 km to 5 km relative

to the Moon geode.

6.4.2 Assumptions

The following assumptions have been taken: • Landing system based on legs • Vertical velocity at Ground Contact as a result of engine cut-off at 1.5 m • Residual horizontal velocity of 1 m/s max. • Landing mass 22 500 kg • Lander attitude error of 2° • Surface slope 10° • Lander geometry:

o Surface Platform Height = 3.65 m (0.65 m Rock/Depression + 0.5 m clearance + 2.5 m Engine Nozzle)

o The Lander shall be able to land over a rock of the given height (this defines the minimum platform clearance to the lunar surface). CoM of lander body above surface platform 6 m.



page 103 of 317

s6.4.3 Lander stability modelling

An analysis has been performed to define the dimension of the lander leg footprint required to maintain a stable platform (no toppling) under the landing conditions specified. The required footprint for stability of lander has been calculated for the following conditions:

• Static Stability: o Worst case approach of slope and one foot or side supported on a rock

• Dynamic Stability: o Assumption Vv = 0 o Worst case assumption Vv = 0, Lander has attitude error o Assumption Vv > 0 o Assumption Vv > 0, Lander has attitude error

For details of the equations used in the model, see Appendix A, Lander stability model.

6.4.4 Lander stability parametric analysis

The static and dynamic stability models have been used to perform a parametric analysis to examine the influence of a number of factors on the minimum required footprint diameter for three-, four- and five-leg systems. The parameters examined are:

• Surface slope change for a constant CoG height • CoG height change for a constant surface slope • Increasing the residual horizontal velocity keeping Vv = 0 m/s • The required minimum footprint diameter for varying Vh & Vv, >0 m/s

Additionally, an estimated system mass has been calculated to assess whether there is an optimum number of legs based upon the mass criteria. For this, the following assumptions have been taken:

• Leg is a tube with internal ø200, external diameter calculated against the maximum allowable yield stress

• Maximum allowable yield stress is factored by 2 • Worst case from bending or buckling has been taken • Aluminium and titanium have been considered • The resulting leg mass is factored up by two to take in to account potential leg supporting

structures Initial input data for the footprint analysis are:

Table 6-10: Input for stability parametric analysis



page 104 of 317

sFigure 6-16 shows the results in graphical form:

Foot-Print Diameter- Static Stability Constraint against increasing Surface Slope, CoG = 7.6 m

0.000 m

5.000 m

10.000 m

15.000 m

20.000 m

25.000 m

30.000 m

35.000 m

5 º 10 º 15 º 20 º 25 º 30 º 35 º 40 º 45 º

Surface Slope [Degs]

Foot

-prin

t dia

met

er [m

]

3 4 5

Figure 6-16: Footprint diameter versus increase in surface slope

Foot-Print Diameter- Static Stability Constraint against increasing CoG Height for Surface Slope 10 deg.

0.000 m

1.000 m

2.000 m

3.000 m

4.000 m

5.000 m

6.000 m

7.000 m

8.000 m

4 5 6 7

CoG Height [m]

Foot

-Prin

t Dia

met

er [m

]

3 4 5

Figure 6-17: Footprint diameter versus increase in CoG height



page 105 of 317

sFoot-Print Diameter- Dynamic Stability Constraint Vv=0 m/s against Horizontal

Velocity

0.000 m

2.000 m

4.000 m

6.000 m

8.000 m

10.000 m

12.000 m

14.000 m

16.000 m

18.000 m

0.5 m/s 1.0 m/s 1.5 m/s 2.0 m/s

Residual Horizontal Velocity [m/s]

Foot

-prin

t dia

met

er [m

]

3 Legs Vv=0 m/s 4 legs Vv=0 m/s 5 Legs Vv=0 m/s

Figure 6-18: Footprint diameter versus increase in residual Vh, Vv=0 m/s

0.0 m/s 0.5 m/s 1.0 m/s 1.5 m/s 2.0 m/s0.0 m/s

0.5 m/s1.0 m/s

1.5 m/s2.0 m/s

0.0 m1.0 m2.0 m3.0 m4.0 m5.0 m6.0 m7.0 m8.0 m9.0 m

10.0 m11.0 m12.0 m13.0 m14.0 m

Foot-print Diameter

Residual Vertical VelocityResidual Horizontal

Velocity

Foot-Print Diameter- Dynamic Stability Constraint at Surface Contact 4 Leg system

13.000 m-14.000 m12.000 m-13.000 m11.000 m-12.000 m10.000 m-11.000 m9.000 m-10.000 m8.000 m-9.000 m7.000 m-8.000 m6.000 m-7.000 m5.000 m-6.000 m4.000 m-5.000 m3.000 m-4.000 m2.000 m-3.000 m1.000 m-2.000 m0.000 m-1.000 m

Figure 6-19: Footprint diameter versus Vh & Vv > 0 m/s

Mass Ratio Estimate.v.Number of LegsVh=Vv=2 m/s, Static Stability Foot-print for CoG= 7 m and 10 deg Slope

0.000

0.200

0.400

0.600

0.800

1.000

1.200

1.400

1.600

1.800

2.000

3.00 4.00 5.00 6.00 7.00

Number of Legs

Rat

io- T

otal

Mas

s E

stim

ate

1200.000 Kg

2200.000 Kg

3200.000 Kg

4200.000 Kg

5200.000 Kg

6200.000 Kg

7200.000 Kg

Al.Alloy- Total Mass Ratio Ti.Alloy- Total Mass Ratio Total mass all legs- al.Alloy Total mass all legs- Ti.Alloy

Figure 6-20: Mass ratio wrt three-leg system for increased number of legs



page 106 of 317

s6.4.5 Leg design

The following results are obtained from the analysis detailed above:

Table 6-11: Parametric analysis input summary

Table 6-12: Parametric analysis result summary

For the four-leg design case chosen, the minimum required footprint, without margin applied, is defined by the static stability case and is 8.095 m. Due to the four-leg configuration, the Lander can in some instances have a cross-axis topple. It can be seen from the data that the cross-axis



page 107 of 317

stopple is 14.535º and remains in a stable three-foot contact configuration if the minimum footprint dimension is used. The estimated leg-loading is given at the bottom of the table. The chosen footprint dimension for the LEV design is 10 m to include margin. Note that the presence of damping elements in the system to damp out the vertical and horizontal velocities causes the footprint to become smaller and therefore as a result increase the instability. The calculation for the selected design case includes this effect in the stability calculation. This results in the final design as shown in Table 6-13:

Table 6-13: Leg design summary

6.5 LEV – configuration 6.5.1 Requirements and design drivers

For the configuration of the LEV, the following set of requirements and design drivers were taken into account:

• The selected launcher, Ariane-5, volumetric capability, which mainly drives the diameter to 4.57 metres.

• Due to the dynamics at landing of the LEV, and the requirements on the landing system (e.g. thrusters, GNC and landing legs) the overall length of the configuration should be kept as low as possible.

• For optimal mass distribution in the descent and ascent phases for the Moon surface missions, the configuration needs to provide a modular design.

• On the surface of the Moon each crewmember should have 12.5 m3 habitable volume.



page 108 of 317

s6.5.2 Assumptions

The requirement for optimal mass distribution in ascent and descent phases results in a configuration design where the LEV consists of three parts: the Descent Module (DM), the Surface Habitation Module (SHM) and the Lunar Ascent Vehicle (LAV). By leaving the SHM with the DM on the surface an optimal AV can be obtained, as shown in Figure 6-21:

DescentModule

Surface Habitation

Module

LunarAscent Vehicle

Figure 6-21 LEV configuration

Due to the limiting diameter of 4.57 metres of the Ariane-5 launcher fairing, the design of the overall height of the LEV resulted in 13.9 metres, as shown in Figure 6-22. :

Figure 6-22 Main LEV dimensions (in mm)



page 109 of 317

sDue to this height, the SHM is launched combined with the AV on top, and the heavy DM is launched separately, and assembly shall take place in LLO, as shown in Figure 6-23:

Figure 6-23: Main elements of the LEV inside the Ariane-5 long fairing

6.5.3 Baseline design

The DM is the unmanned portion of the LEV. To fit in the fairing and to comply with the required footpad distance, a deployable landing system is applied. The stowed DM is shown in Figure 6-30. The overall height is a result of the large main engine nozzle, the required landing leg damping stroke and the large main propellant tank. For the ascent phase, the DM provides a platform for launch.

Figure 6-24: Stowed Descent Module



page 110 of 317

sA truncated cone on the upper part of the DM provides the interface to the SHM structure.

Figure 6-25: Surface Habitation Module with airlock

The SHM houses the crew compartment and is pressurised for a shirtsleeve environment. It is equipped with an airlock; see Figure 6-25. The airlock is accessed by an internal hatch, and provides access to the lunar surface via a large external hatch or door. From the main crew compartment in the SHM, a passage tunnel enables the astronauts to move from the AV to the SHM, providing two separate crew compartments for living and working. The AV is shown in Figure 6-26. This is the module that provides the crew seating during descent and landing as well as the ascent phase. The pressurised compartment includes a small tunnel that provides access to the Hub in orbit around the Moon when docked. The propulsion module has been positioned above the crew cabin to facilitate a stable platform during powered ascent.

Figure 6-26: Ascent Vehicle

The propulsion system needs to be accommodated on the top to show crew passage from the SHM and the AV. This configuration turns out to be favourable as regards GNC also. One issue is the accommodation of a window in the AV for visual guidance capability shortly before landing. Given the selected crew posture in the AV, the only possibility appears to be the



page 111 of 317

sprovision of a ‘rear mirror’ concept that enables visibility even without direct view of the landing site.

6.6 LEV – life support 6.6.1 Requirements and design drivers

Table 6-14 shows the list of requirements applicable to the LEV Life support subsystem design. The main requirements are the number of crew and the mission duration. Those requirements determine the metabolic activity, which is the driver to calculate the consumables (food and oxygen) and the waste (faeces, urine, carbon dioxide and water). The allowance for drinking and hygiene water is of major importance, due to the fact that water represents the highest percentage of the consumables in terms of mass. The number and duration of EVA operations has an impact on both the metabolic activity of the crew and on the H2O and atmosphere losses via airlock and pressure suit. The atmosphere design requirements (volume of pressurised air, pressure and atmosphere composition) determine the atmosphere losses, and have also an impact on the complexity of the system, which depends critically on the pressurised volume to maintain in habitable conditions. The number of modules is also important, since some equipment as fans and pressure equalization valves is needed in the inter-modules. In this case, the two different modules correspond to the Ascent Vehicle Habitation Module, which will be provided of their own life support system each. The consumables will also be calculated for the Ascent and Habitation Module separately.

Number of crew 3 #Mission duration 14 daysDuration of the ascent phase 1 daysTotal number of EVA sorties 12 #Number of crew per EVA 2 #Duration of EVA sortie 6 hoursDrinking water allowance 3.91 kg / CM-dHygiene water allowance 0 kg / CM-dAirlock Volume 4.25 m3

Habitat Volume 40 m3

Return Capsule Volume 7.94 m3

Pressure 48 kPaAtmosphere O2 content 40 %Atmosphere N2 content 60 %Number of Modules 2 #

Mission Design Inputs

Table 6-14: Mission design requirements

6.6.2 Assumptions and trade-offs

6.6.2.1 Crew model

The energy consumption for each of the activities of the crew is shown in Table 6-15 (RD[17]):



page 112 of 317

s

Metabolic Cost of Crew Activity (75kg crew member) kcal Sleep 73 Pre- and post sleep 75 Leisure activities 77 Personal hygiene 150 Eeatin 85 Exercise 250 Station keeping 145 Laboratory activities 150 EVA mission tasks 250 EMU donning/doffing 275 Egress/ingress 225 Pre-EVA setup & post EVA EMU care 200

Table 6-15: Metabolic cost of crew activities (75 kg crewmember)

Three different schedules (see Table 6-16, Figure 6-14 and Figure 6-15) are assumed: EVA day, non-EVA day, and ascent / descent day. An average schedule for all the astronauts is calculated from the mission phases and EVA planning. For the LEV, this average schedule leads to an energy consumption of 3140 kcal / crewmember-day and a heat to be dissipated of 152W / crewmember.

ACTIVITYSleep 8Pre- and post sleep 4Leisure activities 0Personal hygiene 0Eating 2Exercise 0Station keeping 0Laboratory activities 0EVA mission tasks 6EMU donning/doffing 1.5Egress/ingress 1Pre-EVA setup & post EVA EMU care 1.5TOTAL TIME (24hrs) 24

No. of hours

Table 6-16: Schedule for an astronaut on an EVA day

ACTIVITYSleep 8Pre- and post sleep 4Leisure activities 0Personal hygiene 0Eating 2Exercise 0Station keeping 7Laboratory activities 0EVA mission tasks 0EMU donning/doffing 1.5Egress/ingress 0Pre-EVA setup & post EVA EMU care 1.5TOTAL TIME (24hrs) 24

No. of hours

Table 6-17: Schedule for an astronaut on a non-EVA day



page 113 of 317

s ACTIVITYSleep 8Pre- and post sleep 5Leisure

ti iti0

Personal hygiene 1Eating 2Exercise 0Station keeping 8Laboratory

ti iti0

EVA mission t k

0EMU d i /d ffi

0Egress/ingres 0Pre-EVA setup & post EVA EMU 0TOTAL TIME 24

No. of hours

Table 6-18: Schedule for an astronaut on a descent / ascent day

The consumable needs and waste production by the crew is directly related to the energy requirement. The extra assumptions needed are: the diet composition, which is 15% in proteins (average protein composition from RD[18]), 30% fats (palmitic acid) and 55% carbohydrates (glucose); and certain degree of degradation of these components (90, 95 and 98%, respectively). All these data are taken from RD[20]. A mass balance can be then done considering the main elements (C, H, O and N). As a result, the oxygen and dry food requirements, and the urea, faeces, CO2 and metabolic water production rates for the crew are calculated. Table 6-19 shows the stochiometric coefficients of the overall reaction taking place in the average crew organism:

Food Oxygen Faeces CO2 Urea H2OC 1 0 1 1 1 0H 1.9514257 0 1.8891842 0 4 2O 0.6230896 2 0.3943751 2 1 1N 0.0483719 0 0.1103540 0 2 0

1.0718266 1.1658991 0.0469817 1.0015141 0.0233308 0.9547548alpha beta gamma delta epsilon phi

STOECHIMOETRICS (in mols / h)

Stoechio

Table 6-19: Human metabolic activity

These values allow the evaluation of the needs of the crew for the overall mission. The water in food and the drinking water are not taken into account. The production of urea and faeces calculated above allow an estimation of the yellow and black water produced from typical concentrations. The excess of water, including the metabolic water produced, is excreted to the air via skin evaporation:

blackyellowmetabolicDrinkonperspirati WWWWW −−+=

6.6.2.2 Storage

The storage capabilities are calculated from the consumables needed and the products excreted by the crew. Tanks for water, oxygen, and nitrogen are designed taking into account the needs and a 10% margin. The tanks are considered in pairs for safety reasons. Food storage capabilities are calculated from the volumetric coefficient for food including packaging. The value considered is 1.8 l / kg (RD[19]). CO2 removal is done by means of LiOH canisters. The number of canisters needed is also calculated to have at least one extra canister (1 day extra autonomy approximately). A single tank for urine is considered in the habitation module, while in the ascent module, urine and faeces are stored in bags and jettisoned.



page 114 of 317

s6.6.2.3 Atmosphere selection

The atmosphere selection has been done on the basis of a compromise between fire risk and pre-breathing time. An atmosphere with lower pressure will mean lighter structure for the spacecraft. However, fire risk is increased when the oxygen content of the atmosphere is increased. The maximum allowable concentration given a total pressure can be calculated from (RD[21]):

3.101

45.23(%)max2atmP

O =

In addition, when lowering the pressure of the habitat (i.e. from the spacecraft to the pressure suit) the tissue ratio (TR) evaluates the risk of suffering decompression sickness (DCS), TR is calculated as the pressure of inert gas in the tissue (which in equilibrium is equal to the partial pressure of inert gas in the original atmosphere) over the total pressure in the new habitat. In this case, the new habitat is the suit, and nitrogen is the inert gas in the spacecraft atmosphere, thus:

suit

N

Ppp

TR 2=

Americans and Russians consider different values of the TR as acceptable to avoid DCS; TR = 1.6 and 1.8, respectively. However, RD[21] indicates that in the future, due to the more frequent EVA operations, this limit should be reduced to 1.4, value that has been taken into account in this design. Pre-breathing pure O2, is therefore necessary to reduce the partial pressure of inert gas in the tissue. The decay is modelled by an exponential curve with certain half decay constant. This constant, with the current device (an O2 mask), takes the value of 6 hours.

tkNN epptpp ⋅−⋅= )0()(

22

The above equations, together with an acceptable limit for TR, allow calculating the O2 content of the atmosphere to have no pre-breathing (or a given pre-breathing time). The total pressure of the suit needs also to be selected. A soft suit of Apollo type (26 kPa) has been selected due to various factors, including mobility and weight. Figure 6-27 shows the atmosphere oxygen content versus the total pressure. The blue line represents the optimum atmosphere composition for no pre-breathing, and the red line the optimum atmosphere to not increase the fire risk. The areas in gray are the acceptable atmosphere compositions for human life. It can be seen that the optimum atmosphere falls under the irrespirable area, which cannot be accepted. This means that either an increase of fire risk has to be considered, or a certain pre-breathing time needs to be accounted for.



page 115 of 317

s

Figure 6-27: Atmosphere composition

Figure 6-28 shows the final atmosphere selection, which is a compromise between both limitations. The blue line now represents 60 minutes pre-breathing time and the red line represents a 15% higher O2 concentration than allowed by the first equation of this section.

Figure 6-28: Atmosphere composition (continued)

The final atmosphere selection for the LEV is the cross point, which is at 40% O2 / 60% N2 with a total pressure of 48 kPa.

6.6.2.4 Cabin / atmosphere losses model

There is an extra need for O2 and N2 due to the fact that the atmosphere has to be delivered and made up for the cabin losses. The atmosphere losses are evaluated from ISS module specification, which is 83 kg /year (RD[22]). Leakage is modeled with this value as a reference, taking into account that the internal pressure is the driving force



page 116 of 317

s)(_83

3.1012

2ydurationMission

ppLeakage O

O ⋅⋅=

)(_833.1012

2ydurationMission

ppLeakage N

N ⋅⋅=

6.6.2.5 EVA operations

EVA operations affect the consumable design in two ways. First, EVA results in an increase of the energy expenditure by the crew, secondly, there are losses associated to EVA operations that need to be taken into account. These are:

• H2O losses for thermal control: the value taken is 0.57 kg / CM-h from RD[22] • O2 losses from the suit: 0.005 kg / CM-h from RD[22]

Atmosphere losses from Airlock Operations: given the size of the airlock (4.25 m3 for two crewmembers) and size of the suited crewmembers (0.28 m3 / CM), part of the atmosphere is lost in each EVA sortie. Using a pump to save part of the atmosphere before opening the hatch makes sense only for very large airlocks and / or multiple EVA operations, which is not the case for the LEV mission.


The type of life support system selected for the LEV is an open loop life support system, due to the short duration of the mission. Figure 6-29 shows the schematic of the Life Support System of the surface Habitation Module.

Figure 6-29: LEV descent module life support system

The Life support system of the Ascent Vehicle is very similar. The main differences are the sizes of tanks and storage capacity, and the absence of airlock, water heater, urinal and major constituents analyser. The relevant schematic is shown in Figure 6-30.



page 117 of 317

s

Figure 6-30: LEV ascent module life support system schematic

6.6.4 List of equipment

Table 6-20 shows the components considered for the design of the Ascent Vehicle:

Element 3 Unit Name1 Heat Exchanger 1 19.9 Fully developed2 Water Separator 2 9.8 Fully developed3 Fan and motor 2 18.6 Fully developed4 Oxygen Valve 9 2.0 Fully developed5 Nitrogen Valve 9 2.0 Fully developed6 Water Valve 13 1.8 Fully developed7 Water Pressure Regulator 2 0.4 Fully developed8 Oxygen Pressure Regulator 8 0.9 Fully developed9 Nitrogen Pressure Regulator 8 0.9 Fully developed10 LiOH Cartridge 2 17.4 Fully developed11 ISS HEPA Bacteria Filter Assembly 1 4.9 Fully developed13 Transducer and switch 4 0.5 Fully developed14 Fluid interface connection 80 0.5 To be modified15 Pump and motor 1 24.0 Fully developed16 Filter 10 0.1 To be developed20 Cabin Pressure Relief Valve 2 5.4 Fully developed21 Potable Water Gun 1 5.0 To be developed24 Personal Hygiene kit 3 1.5 To be developed25 Medical Kit Ascent 1 14.1 Fully developed27 Smoke detectors 1 1.5 Fully developed28 Portable Fire Extinguishers 2 7.8 Fully developed29 Apollo Suit 3 15.5 Fully developed32 Ascent Water tank 2 1.5 To be modified35 Ascent Oxygen Tank 2 0.9 To be modified36 Ascent Nitrogen Tank 2 0.2 To be modified38 LEB Food Storage Box Ascent 1 5.0 Fully developed40 HEPA Filter 2 2.0 Fully developed42 LIOH Canister Ascent 2 6.0 Fully developed

28 381.6ELEMENT 3 SUBSYSTEM TOTAL

Element 3: Lunar Ascent Vehicle MASS [kg]Unit Quantity Mass per Maturity Level

Table 6-20: LEV life support equipment summary



page 118 of 317

sTable 6-21 shows the items and quantities that have been considered for the design of the Surface Habitation Module:

Element 2 Unit Name1 Heat Exchanger 3 19.9 Fully developed2 Water Separator 2 9.8 Fully developed3 Fan and motor 2 18.6 Fully developed4 Oxygen Valve 9 2.0 Fully developed5 Nitrogen Valve 9 2.0 Fully developed6 Water Valve 15 1.8 Fully developed7 Water Pressure Regulator 2 0.4 Fully developed8 Oxygen Pressure Regulator 8 0.9 Fully developed9 Nitrogen Pressure Regulator 8 0.9 Fully developed10 LiOH Cartridge 2 17.4 Fully developed11 ISS HEPA Bacteria Filter Assembly 4 4.9 Fully developed12 MCA (Major Constituents Analyzer) 1 54.7 To be modified13 Transducer and switch 4 0.5 Fully developed14 Fluid interface connection 400 0.5 To be modified15 Pump and motor 1 24.0 Fully developed16 Filter 10 0.1 To be developed17 IMV Valve 1 5.1 Fully developed18 IMV Fan 1 4.2 Fully developed19 Manual Pressure Equalization Valve 2 1.1 Fully developed20 Cabin Pressure Relief Valve 4 5.4 Fully developed21 Potable Water Gun 1 5.0 To be developed22 Potable Water Heater 1 5.0 To be developed23 Commode/Urinal 1 50.0 To be developed26 Medical Kit Descent 1 89.5 Fully developed27 Smoke detectors 3 1.5 Fully developed28 Portable Fire Extinguishers 2 7.8 Fully developed29 Apollo Suit 3 15.5 Fully developed30 Apollo PLSS 3 75.2 Fully developed31 Descent Water Tank 2 31.0 To be modified33 Descent Oxygen Tank 2 21.4 To be modified34 Descent Nitrogen Tank 2 15.1 To be modified37 LEB Food Storage Box Descent 4 5.0 Fully developed39 Urine Storage Tank 1 17.3 To be modified40 HEPA Filter 8 2.0 Fully developed41 LIOH Canister Descent 14 6.0 Fully developed

35 1277.3ELEMENT 2 SUBSYSTEM TOTAL

Element 2: Surface Habitation ModuleUnit Quantity Maturity LevelMass per

MASS [kg]

Table 6-21: SHM life support equipment summary

6.6.5 Budgets

The equipment used has been selected by choosing the optimum components from a database including ISS, Shuttle, Spacelab, CRV, Apollo and Soyuz data, The final life support system mass and power budget is shown in Table 6-22 and Table 6-23.



page 119 of 317

sTotal Life Support Mass 401.49 kg- Total Equipment Mass 381.64 kg Margin 6.0%- Total Consumables Mass 19.79 kg Oxygen 3.44 kg Potable Water 12.90 kg Hygiene Water 0.00 kg Food (dry) 1.90 kg Food Packaging 0.81 kg Nitrogen 0.73 kgTotal Power 342.94 W

Output Ascent Vehicle

Table 6-22: Mass and power budgets of ascent vehicle life support system

Total Life Support Mass 2108.85 kg- Total Equipment Mass 1658.91 kg Margin 7.0%- Total Consumables Mass 449.87 kg Oxygen 83.61 kg Potable Water 270.93 kg Hygiene Water 0.00 kg Food (dry) 26.58 kg Food Packaging 11.34 kg Nitrogen 57.42 kgTotal Power 978.51 W

Output Descent Vehicle

Table 6-23: Mass and power budgets of surface habitation module life support system

6.7 LEV – propulsion This section describes the ascent and descent propulsion system of the Lunar Excursion Vehicle (LEV). The purpose of the LEV is to bring astronauts to the surface of the Moon from Low Lunar Orbit (LLO) and back. Both the ascent and descent propulsion stages of the LEV have many similarities with the design of the propulsion system of the Apollo LEM (Lunar Excursion Module). The ascent and descent propulsion systems use storable propellants. Both bi-propellant systems use MON (mixed oxidisers of nitrogen) as oxidiser and MMH (mono methyl hydrazine) as fuel.



page 120 of 317

s

Figure 6-31: RS-41 engine and vehicle configuration

The ascent propulsion system (APS) comprises four RS-41 (US, Rocketdyne) engines; (see Figure 6-31) each capable of producing 12 kN of thrust in vacuum conditions. The RS-41 is pressure fed using either nitrogen or helium and hence the system reliability is maximised compared to other feeding systems e.g. pump-fed systems. The descent propulsion system (DPS) uses a single throttable engine to meet two major requirements: hovering capability and throttability. Note that no throttable storable engines in the defined thrust range exist today so, this design is based on technology extrapolation. The DPS is required to meet the following major requirements:

• 10:1 throttability (140 kN to 14 kN) and gimballing • 140 kN is the maximum design thrust level • Hovering capability (between 83 kN and 39 kN)

Most existing engines (basically launcher engines) are either out of production, have a low burn time, are not restartable or have low performance. Vehicle hovering (throttability) can be achieved using the following methods:

• Pump-fed system (not available today) • A dedicated throttable pressure fed engine is developed, as in the Apollo programme • A sufficient number of discretised thrust levels using numerous engines

A sophisticated pump-fed engine is considered very complex and has a higher failure rate compared to the other options mentioned above. Therefore, for simplicity, this option is discarded in this study. Throttability can also be achieved by using a number of smaller thrusters/engines in a cluster. However, this way of discretising the thrust is limited and hovering cannot be achieved anymore if the design mass increases. Hence, this option is discarded. As a result of the discussion above, a new development of a DPS engine is required. A single pressure fed engine seems to be the best option for the descent stage in terms of reliability and



page 121 of 317

sperformance. This choice is also based on the Apollo DPS design. Moreover, the ability to tailor the engine for this particular mission is an advantage.

6.7.1 Requirements and design drivers

The APS is required to perform a velocity increment of ∆V = 2100 m/s to achieve LLO from the lunar surface. The total propellant mass required for the ascent is 5 044 kg (3 141 kg of MON and 1 903 kg of MMH). This sizes both the propellant tanks and the pressurant tanks. The DPS is required to perform a velocity increment of ∆V = 2300 m/s using storable propellant to descent from LLO to lunar surface. The DPS is required to be throttable and to produce the necessary thrust in the LEV hovering phase. The total propellant mass required for the ascent phase is 23437 kg (14593 kg of MON and 8844 kg of MMH).


A propulsion system that uses a pump-fed engine is usually lighter compared to a propulsion system using a pressure-fed engine which requires higher tank pressure. It is well understood that the mass penalty of using a pressure-fed engine reduces the payload performance of the vehicle. However, a pressure-fed engine does not incorporate the same complexity as a pump-fed engine i.e. pumps and turbines. Taken this into consideration the reliability is higher for a pressure-fed propulsion system compared to pump-fed systems.

Figure 6-32: DPS engine cluster configuration

A configuration trade-off was performed for the DPS to evaluate whether it is plausible to obtain hovering capability using a mixed cluster of small and large engines. Many options were studied and the most promising one is presented here. As a baseline, the Apollo descent trajectory is used. By combining four Aestus engines and 13 R-40B engines (see Figure 6-33), it is possible to ‘mimic’ the Apollo descent trajectory quite well:



page 122 of 317

s

Figure 6-33: Descent trajectory thrust profile

Figure 6-33 shows the Apollo descent trajectory and the thrust profile achieved by using four Aestus engines and 13 R-40B engines. This engine combination can produce the desired thrust in the LEV descent phase. However, a cluster of engines can only produce a discretised thrust profile. Therefore, if the total mass of the vehicle increases or decreases during the design phase it is not certain that the engine combination presented above can still perform the task. The hovering capability using a cluster of engines is therefore very sensitive to mass changes. For this reason this approach has been rated low.


6.7.3.1 Ascent propulsion system

The APS baseline design is a pressure-fed propulsion system comprising four RS-41 main engines and 16 X 500 N EAM for AOC (Attitude and Orbit Control). The APS is a combined propulsion system (CPS) meaning that the AOCS thrusters and main engines feed from the same propellant tanks. The bi-propellant system uses MON and MMH with a nominal mixture ratio of 1.65 to maximise performance and to equalise oxidiser and fuel tank volumes.



page 123 of 317

s

Figure 6-34: APS architecture

Figure 6-34 shows the APS schematic and architecture with two oxidiser and two fuel tanks. The propellants tanks are all metal, same dimensions as on the ATV. The domes are spin formed using Ti-15-3 and the minimum wall thickness is 1.2 mm. The maximum expected operating tank pressure (MEOP) is 25 bar.

Figure 6-35: APS propellant tank

Figure 6-35 shows the equatorial ring on the propellant tank used to attach the tanks to the vehicle structure. The propellant tank mass incl. PMD is approximately 80 kg and the outer diameter is 1.28 m.



page 124 of 317

sThe APS comprises two composite over-wrapped pressure vessels (COPV) to pressurise the propellant tanks during the mission. The COPV has a thin inner titanium liner and is over-wrapped with graphite epoxy layer that enables a considerable mass reduction. The pressurant tanks are pressurised to 250 bars at Beginning Of Life (BOL). The APS has full RCT redundancy and there is no mission impact if a single RCT failure occurs. Pyrotechnic valves and pressure regulators have all internal redundancies to enhance reliability. All components have demonstrated long-term compatibility with propellants.

6.7.3.2 Descent propulsion system

Figure 6-36: DPS engine configuration

The DPS comprises a single main engine capable of hovering the LEV (throttable). The DPS main engine must be capable of throttling through a 10:1 thrust range, be gimballed, and be able to hover at any time during the descent phase. The DPS configuration is shown in Figure 6-36. The DPS is a bi-propellant pressure-fed system using MON and MMH at a nominal mixture ratio of 1.65 for maximum performance. The propellant is stored in one titanium-alloy common bulkhead storable-liquid propellant tank with a diameter of 3.52 m and a total length of 3.26 m.



page 125 of 317

s6.7.4 List of equipment

6.7.4.1 Ascent propulsion system

Table 6-24 shows the ratio of dry mass to wet mass for the LEV APS. Here the dry mass includes the pressurant mass, as well as a 10% subsystem margin. A 20% system margin is also included as part of this design.

Value Unit

Dry Mass 871 kgPropellant 504 kgTotal Wet Mass 591 kg

Table 6-24: LEV APS total mass budget

All masses are calculated using the characteristics of existing components or using the actual mass of existing components where possible.

Component Qty Mass Main 4 69 Oxidiser 2 80 Fuel 2 80 Pipework & Valves 1 44 Pressurant Tank 2 28 Reaction Control 16 5 Pressurant 1 15

Table 6-25: LEV APS equipment summary and dry mass budget

Table 6-25 shows a mass budget of the LEV APS.

6.7.4.2 Descent propulsion system

Table 6-26 shows the ratio of dry mass to wet mass for the LEV DPS. Here the dry mass includes the pressurant mass, as well as a 10% subsystem margin. A 20% system margin is also included as part of this design.

Value UnitDry Mass 1859 kgPropellant Mass 23437 kgTotal Wet Mass 25296 kg Table 6-26: LEV DPS total mass budget



page 126 of 317

sComponent Qty Mass (kg)

Main Engine 1 300Common Bulkhead Tank 1 818Pressurant 1 65Pipework 1 69Pressurant tank 2 112Gimbal 1 204

Table 6-27: LEV DPS dry mass budget

Table 6-27 shows the LEV DPS dry mass budget (excl. margin). The main engine has a design mass of 300 kg and estimated on the basis of current engine technology within this thrust range.

6.8 LEV – GNC 6.8.1 Introduction

The lunar excursion vehicle (LEV) is made of three main subassemblies: 1. The Lunar Ascent Vehicle (AV) 2. The Surface Habitation Module (SHM) 3. The Descent Module (DM)

The flight controls and displays of the LEV are located on the AV. The ascent engine systems are also installed on the AV. The SHM will be used only during surface operations. The DM houses the landing legs and descent engine system and its propellant. The AV and the SHM are launched together as a major subassembly (S/A). The lunar orbit injection (LOI) propulsion module (PM) is also part of the S/A. The two trans-lunar injection (TLI) PMs are launched separately. The TLI PMs dock with the AV+SHM S/A in low Earth orbit (LEO). The DM is launched on its own. Its LOI PM and TLI PMs are launched separately. At the time of the AV+SHM S/A and DM TLIs the Hub is already in low lunar orbit (LLO). Once in LLO the AV+SHM S/A docks with the Hub and then the DM docks with AV+SHM already at the Hub. The modules do not have a crew on board from launch to commencement of lunar exploration mission.


The following requirements for the GNC subsystem of the LEV have been identified concerning assembly:

• The PMs shall autonomously dock with the respective subassembly (S/As)1 • The AV+SHM S/A shall autonomously dock with the Hub • The DM shall autonomously dock with the AV+SHM S/A already at the Hub

1 From here on the term S/A refers to the AV+SHM S/A and to the DM interchangeably.



page 127 of 317

sThe requirements used for the design of the LEV GNC system during the descent part of the mission are:

• The LEV shall land safely on the Moon from a polar elliptic orbit of 100 km altitude • The horizontal velocity with respect to the surface shall be less than 0.5 m/s at an altitude

of 1.5 m • The vertical velocity shall be null at an altitude of 1.5 m • The GNC system shall tolerate a shift in the CoM of 1.4 m • The GNC system shall tolerate a change in the pitch of 90˚ • The GNC system shall allow manual takeover of the landing from an altitude of 250 m • The GNC system shall provide the capability to perform an autonomous abort to Hub

orbit from any point of the descent trajectory The requirements for the AV GNC system on the ascent part of the mission are:

• The GNC system shall be able to autonomously ascend, acquire the orbit, and rendezvous and dock (RVD) with the Hub

• The GNC shall allow manual takeover of the RVD manoeuvres from a distance of 100 m • The GNC shall provide collision avoidance guidance and control at any point of the RVD

trajectory • The collision avoidance mode shall be triggered either manually or automatically

Other general requirements are:

• The GNC system should allow an operator to take over the control of the S/C or at least command an abort during any phase of the mission. The operator can be the crew on board, e.g., during the LEV descent, or remotely located, e.g. on the Hub or on Earth, e.g., during RVS manoeuvres

• The sensor systems used should be highly reliable and readily integrated in an autonomous GNC system

Current and past European planetary landing efforts focused on robotic missions. The expertise obtained in this area should be employed to the maximum. However, extending the GNC algorithm work to a manned mission is not immediate and development, implementation, integration, and testing of the GNC algorithms should be pursued as early as possible Note that the above requirements and design drivers are not exhaustive so that they should be treated as a set of guidelines for future studies rather than a frozen set of requirements. For example the altitude of 250 m from where the crew can take over the control of the landing manoeuvre is chosen based on what is believed to be a good compromise between efficient propellant use and visibility of the landing site.


It has been assumed that all LEV operations, from assembly in LEO to lunar landing, are performed autonomously. As described in the previous sections the crew is able to command an abort or take over the control of the manoeuvres during all mission phases except during the descent orbit injection (DOI) burn. However, the crew can postpone or abort the descent either before the DOI, or abort immediately after the DOI.



page 128 of 317

sIt has been assumed that only thrusters are used as actuators of the LEV GNC system. It has been assumed that the orbit phasing for the LEO assembly operations is performed under command from the ground station (GS).


The following paragraphs describe the GNC systems according to the sequence of the mission phases. The GNC system has been designed for autonomy.

6.8.4.1 LEO assembly

The LEO assembly consists of autonomous RVD operations between the major S/As and the respective Propulsion Modules. The LEO assembly is similar for both AV+SHM S/A and the DM. The only difference is that the AV+SHM S/A has two PMs to be docked with and the DM has three. The PMs are the interceptors and the S/A is the target. One PM at a time will dock with the S/A. During this part of the mission the GNC architecture is centralised. The architecture is shown in Figure 6-37. The PMs have a relatively simple GNC system.

Figure 6-37: LEO assembly GNC architecture

The sensing, processing, and command generation are performed on the S/As (AV+SHM S/A or DM). The commands are sent to the PMs via the ISL. The PM GNC algorithms interpret the commands and the PM onboard computer (OBC) sends the thruster firing commands to the thruster block. The justification for the centralised architecture is the fact that the GNC systems sensors and computer on the S/As need to be used also for RVD with the Hub and the PMs will be instead ejected. The following sensors are employed for the LEO assembly operations:

• Differential GPS (DGPS) o Range: 1 km < R < ~10 km o Requires a GPS receiver on each element and an ISL



page 129 of 317

s• Radiofrequency navigation (RFN)

o Range: 1 km < R < 4 km o Requires Tx and Rx on both elements (S/A and PM.) o Has an ISL capability at transmission rates of up to 9 kbps

• Rendezvous sensor (RVS): o Range: 0 < R < 1 km o Has a laser scanning head and a computer box on the S/A and retroreflectors (RR) on

the PMs. The ranges at which the sensors are used during LEO assembly are shown in Figure 6-38. It is natural to define two safety spheres, based on the ranges of the RFN and RVS. The outer sphere has a radius of 4 km and the range of the inner sphere has a radius of 1 km. Inside the safety spheres, two sensors are used simultaneously. Between the outer and the inner sphere the DGPS and RFN systems provide the dual redundancy and inside the inner sphere the dual redundancy is provided by the RFN and the RVS. The dual redundancy is shown in Figure 6-38. If a significant drift between the estimates of the two sensors is detected, the PM is commanded to go into a safe hold or collision avoidance mode. The PMs and S/As employ autonomous star trackers for estimation of their inertial attitude. The DGPS, RFN, and RVS sensors have the capability to determine the relative attitude. The following sequence is proposed for the RVS of the PMs with the S/As. The sequencing corresponds to the range of the relative navigation sensors employed:

1. The PMs are brought within a few tens of kilometres of the S/A by the GS 2. The DGPS is initialised and the command authority is passed to the S/A 3. The S/A commands the PM to approach 4. Once the outer safety sphere is reached the PM holds its position while the RFN system

is brought online 5. Once the RFN is online the PM proceeds with approaching the S/A 6. When the range of the RVS is reached the PM holds its position again and RVS is

brought online and the DGPS is brought offline 7. Once the RVS is online the PM proceeds with approaching the S/A to docking 8. The PM is commanded to hold at the edge of the outer safety sphere and the RFN system

is brought online 9. Once the RFN is online the PM is commanded to approach the S/A until it reaches the

edge of the inner safety sphere 10. The PM is commanded to hold at the edge of the inner safety sphere and the RVS is

brought online 11. Once the RVS is online, the PM is commanded to approach and dock with the S/A



page 130 of 317

s

Figure 6-38: Range of the sensors employed in the LEO assembly operations

The radius of the outer safety sphere of 4 km is given by the range of the RFN subsystem. This is the range currently specified for the RFN system developed for the Darwin mission. A sketch of the final phase of the RVD between the PM and the AV+SHM S/A is shown in Figure 6-39. The sketch shows the PMs assembled in parallel but the concept proposed also works for the configuration where the PMs are stacked. The RRs would have to be installed on an appendage visible to the RVS head on the S/As.

Figure 6-39: TLIPM docking with the AV+SHM S/A, showing the RVS operation

6.8.4.2 Rendezvous and docking with the Hub

This section presents the RVS sequences and GNC architecture for the docking of the AV+SHM and the DM with the Hub. The design approach for the RVD of the AV+SHM S/A with the Hub, in LLO, is similar to that of the PMs. Three sensosr are employed to obtain relative navigation estimates. The major difference is that the estimation, processing, and commanding is performed on the manoeuvring vehicle, i.e., on the AV+SHM S/A. Another difference is that the DGPS is now replaced with a multimode Doppler radar (MMDR). The following sensors are employed for the LLO assembly operations:



page 131 of 317

s• Multimode Doppler radar (MMDR)

o Range: R > 10 km o Requires a XPDR (XPNDR) on the Hub

• Radiofrequency navigation (RFN) o Range: 1 km < R < 4 km o Requires Tx and Rx on both elements (AV+SHM and Hub) o Has an ISL capability at transmission rates of up to 9 kbps

• Rendezvous sensor (RVS) o Range: 0 < R < 1 km o Has a laser scanning head and a computer box on the S/A and RRs on the PMs

The ranges at which the sensors are used during LLO RVD are shown in Figure 6-40:

Figure 6-40: Range of the sensors employed in the lunar orbit RVD

The only difference between this set of sensors and that of the LEO assembly (Figure 6-38) is that the MMDR, instead of the DGPS, is employed for long-range relative navigation. As an additional safety feature, a vision-based autonomous docking system can be employed to provide triple redundancy during the proximity operations (R < 1 km.). The manoeuvre sequencing is similar to that of the LEO assembly. If a significant drift between the estimates of any of the two sensors is detected the AV+SHM S/A goes into a safe hold or collision avoidance mode. Figure 6-41 shows the final phase of the RVD between the AV+SHM S/A and the Hub. The RFN and RVS systems are used at short range, R < 1 km.



page 132 of 317

s

Figure 6-41: Final phase of the AV+SHM docking with the Hub

Both the RFN and the RVS systems are used. The MMDR is not shown. The AV+SHM S/A performs an autonomous docking with the Hub. The DM docks with the AV+SHM S/A at the Hub. The approach is similar to that of the RVD manoeuvre of the PMs RVD in to the S/As in LEO, and is shown in Figure 6-37. All the sensing is performed at the AV+SHM. The LA+SHM GNC system generates the commands for the DM and sends them to the DM over the ISL. Figure 6-42 shows the final phase of the RVD between the DM and the AV+SHM S/A. The RFN and RVS systems are used at short range, R < 1 km. Both the RFN and the RVS systems are used. The commands are sent from AV+SHM S/A to the DM through the ILS.

Figure 6-42: Final phase of the DM docking with the AV+SHM S/A



page 133 of 317

s6.8.4.2.1 Summary

Table 6-28 shows the sensors used for relative navigation during the unmanned phases of the mission, that is, assembly in LEO and RVD in LLO. There are three major types of GNC algorithms:

1. One algorithm is used on a centralised architecture such as that shown in Figure 6-37. The algorithm implements GNC functions which command an interceptor based on relative navigation estimates obtained by the target. The algorithm runs on the target OBC. This algorithm is applied to the RVD in LEO of the PMs with the AV+SHM S/A and with the DM, and to the DM RVD in LLO with the AV+SHM S/A at the Hub.

2. The second algorithm type is a typical autonomous RVD algorithm. It implements the GNC functions which command the interceptor based on the relative navigation estimates obtained onboard the interceptor. This algorithm is applied to the RVD of the AV+SHM S/A to the Hub.

3. The algorithm running on an interceptor commanded by its target. This is the case of the PM in LEO and of the DM in LLO.

Throughout the mission the inertial attitudes of the S/C are determined by ASTs. In addition to the sensors described above, each of the spacecraft has a set of coarse Sun sensors installed so that they provide full sky coverage. The Sun sensors are employed to determine a departure from a commanded attitude. If such a departure is detected, an emergency Sun acquisition manoeuvre is triggered.



page 134 of 317

s

Range

Module Phase Long (> 10 km)

Medium (~4 km)

Short (< 1 km)

GNC alg. type

Observations

PM LEO DGPS RFN RR III PM is an interceptor commanded by the target.

LEO DGPS I AV+SHM is the target commanding the PM interceptor

RVD to Hub II Autonomous RVD AV+SHM

LLO DM

RVD

MMDR RFN RDS

I AV+SHM is the target commanding the DM interceptor

LEO DGPS RDS I DM is the target commanding the PM interceptors DM

LLO Transponder

RFN

RR III DM is an interceptor commanded by the target

Hub LLO Transponder RFN RR N/A Hub is a passive target

Table 6-28: Summary of the sensors used in LEO and lunar orbit operations

6.8.4.3 Descent from Low Lunar Orbit to lunar surface

The phases of the descent part of the mission are presented below. Figure 6-43 shows the main points of the descent phases. DOI stands for descent orbit injection, PDI for powered descent initiation and TD for touchdown.

Figure 6-43: Points of the descent phases

The descent phases are:

1. Separation to DOI: During this phase of the descent, the navigation is provided by the inertial measurement unit (IMU). The attitude with respect to inertial reference is provided by the AST.



page 135 of 317

s• The DOI is commanded by the on-board computer. The crew can override the

initiation of the DOI burn by either cancelling or postponing it. • The DOI burn puts the LEV into an elliptic orbit with the periapsis at 15 km

altitude. 2. DOI to powered descent initiation (PDI):

• During this phase of the descent, the navigation is provided by the IMU. The attitude with respect to inertial reference is provided by the AST. Close to PDI the MMDR is brought on line and the velocity and altitude estimates of the MMDR are used to update the IMU.

• The initialisation of the PDI is commanded by the on-board computer. The crew can only abort the mission. The crew cannot modify the initialisation of the PDI.

3. PDI to touchdown (TD): This phase has been divided into three subphases: • 1 km<H<15 km. During this subphase, the LEV navigation is provided by the

MMDR which updates the IMU. The MMDR tracks the primary landing site and a few alternate landing sites simultaneously. An illustration of this subphase of the descent is shown in Figure 6-44.

• 25 m<H<1 km. During this subphase, the LEV navigation is provided by the MMDR and a vision or lidar-based system. The vision or lidar system has the capability to detect and avoid obstacles. Obstacle avoidance is implemented by real-time modification of the guidance profile. From the moment when the motion has a predominantly vertical component the MMDR will not be able to estimate the horizontal velocity accurately. As a consequence the MMDR will be used as a radar altimeter only and the vision or lidar system will provide horizontal navigation data.

• 0<H<25 m. During this subphase, the navigation is handed over to the IMU. It is estimated that the vision or lidar system will be blinded by the dust lifted by the descent engine.

Throughout the phase the attitude is provided by an AST. Preliminary trajectory design estimates that the duration of this phase is about 550 s.

Figure 6-44: High-altitude subphase of the powered descent

During the descent from LLO to the surface, the crew is on board the LEV. The GNC system is designed so that a fully automatic landing can be accomplished. The crew is able to command an abort or take over the control of the manoeuvres during all descent phases except during the DOI



page 136 of 317

sburn. However, the crew can postpone or abort the descent either before the DOI, or abort the descent immediately after the DOI all the way to TD. The crew can take over the control of the LEV and manually manoeuvre it from an altitude determined based on tests. The crew interface (C I/F) is made of computer displays, controllers, i.e. joysticks, and various switches needed to switch on and off the LEV subsystems. The vision or laser system is employed for autonomous hazard avoidance down to an altitude where it is blinded by dust.

Figure 6-45: Low-altitude subphase of the powered descent

A preliminary estimate of the control torques required during the descent has been performed. The descent manoeuvre pitch profile has been obtained from the baseline trajectory and is shown in Figure 6-46. Note that the pitch variation is rather mild. The maximum pitch acceleration of 0.026 deg/s2 is encountered at the end of the phase, prior to touchdown. The moment of inertia of the LEV about the transversal axis has been estimated to be Iyy=Izz=382×10 3kgm2.at the beginning of the powered descent phase and Iyy=Izz=158.4×103 kgm2 at the end of the phase. The assumptions made were that the initial mass of the LEV is 51 tonnes, the final mass is 24 tonnes and that the LEV is a cylinder of 10 m length and 4 m diameter, with a uniform mass distribution. It also has been assumed that the moment of inertia varies linearly with time. The required torque during the powered descent phase is shown in Figure 6-47. The maximum torque required is T=74 Nm. It is estimated that the 500 N attitude control thrusters installed on both the AV and the DM will have the capability to provide the required torques. Note that the axial position of its centre of mass (CoM) will shift by 1.64 m towards the AV. The GNC algorithms are to be designed to cope with this relatively large shift of the CoM.



page 137 of 317

s

Figure 6-46: Pitch profile, rate, and acceleration during the descent part of the mission



page 138 of 317

s

Figure 6-47: Time variation of required torque to pitch LEV during powered descent phase

6.8.4.3.1 Summary

Subphase Sensor Observations

Separation to DOI IMU Navigation based solely on IMU

DOI to PDI IMU IMU with the possibility of MMDR updates toward the end of the subphase

High altitude (~15 km) MMDR

MMDR used for navigation towards the target and to update the IMU.

PDI to TD Medium altitude

(a few km)

MMDR + optical

The optical navigation system modifies the guidance profile employed to avoid hazards. When the motion become predominantly vertical, the MMDR is used as a radar altimeter only.

Low altitude (< 1 km)

Optical + IMU

The optical system provides the navigation down to the last tens of metres. The IMU only is employed during the last tens of metres.

Table 6-29: Sensors employed during descent from LLO to lunar surface

The optical sensor denotes either a vision or lidar sensor.

6.8.4.4 Ascent from lunar surface to Low Lunar Orbit

Of the entire LEV, only the AV will lift off and dock to the Hub. The MMDR is employed as a tracking radar to update the state of the IMU prior to launch. An XPDR will be installed on the Hub. The MMDR will provide updates of the orbit of the Hub. Once in orbit, the sequence is



page 139 of 317

ssimilar to that of RVD docking when coming from Earth. Figure 6-48 shows the phases of the ascent. At AV lift off (a), the MMDR is employed to obtain an estimate of the Hub’s orbit.

Figure 6-48: Phases of the ascent part of the mission


The GNC list of equipment is shown in Table 6-30. Mass and power are per unit. C I/F stands for crew interfaces. Only the sensors needed for LEV operations are shown here.



page 140 of 317

s

Equipment Module

Name No. of units Mass (kg) Power

(W) Observations

RR 3 0.25 N/A CSS 6 0.10 0.1 AST 2 5.00 20.0 RFN 1 6.00 12.0

PM

DGPS 1 5.00 10.0 RR 3 0.25 N/A CSS 6 0.10 0.1 AST 2 5.00 20.0 RVS 2H+1C 5.00 10.00 45.0 Two sensor heads one computer. RFN 1 6.00 12.0

DGPS 1 5.00 10.0 MMDR 1 70.00 50.0 ONHA 1 3.00 10.0 C I/F 2 20.00 100.0

AV

IMU 1 3.000 15.0 RR 3 0.25 N/A CSS 6 0.10 0.1 RVS 1H+1C 5.00 10.00 45.0 One sensor head one computer. RFN 1 6.00 12.0

DGPS 1 5.00 10.0

DM

XPDR 1 1.00 5.0 Requires 200W power at the antenna. RR 3 0.25 N/A Hub

XPDR 1 1.00 50.0 Requires 200W power at the antenna.

Table 6-30: AOCS equipment list

6.9 LEV – structures 6.9.1 Requirements and design drivers

For the structure of the Lunar Exploration Vehicle (LEV) the following set of requirements and design drivers were taken into account:

• Structural design shall aim for simple load paths, maximise the use of conventional materials, and simplify interfaces and easy integration.

• The structure shall be designed to meet the requirements for stiffness under the specified load and boundary conditions, this is to avoid dynamic coupling between the low-frequency launcher and LEV modes.

• The structure shall be of adequate strength to withstand the design loads (ground and test loads, launch loads and in-orbit loads) without yielding, failing or exhibiting excessive deformations that can endanger the mission objectives.

• The structure shall be protected from micrometeoroid and debris impact to prevent the risk of catastrophic failures.



page 141 of 317

s6.9.2 Assumptions

The SHM is launched with the AV attached on top; the DM is launched separately (see Figure 6-49):

Figure 6-49: Main elements of the lunar exploration vehicle inside the Ariane-5 long fairing

For the stiffness requirement, the following is assumed: the fundamental frequency for the SHM/AV, and the DM, hard mounted at the separation plane with the launcher, is greater than 9 Hz in lateral direction and greater than 27 Hz in longitudinal direction. As regards strength, the following load cases are assumed to be dimensioning:

• Compression loads from axial and moment loads during launch • Landing loads on the lunar surface • Stresses from internal pressure in the AV and SHM

Multiple values of factors of safety (FS) are used for structures to reflect the varying confidence in different structures, or increased conservatism when the vehicle is manned. During launch and for the landing loads, an ultimate FS of 1.4 is assumed and 1.1 at yield. For the internal pressure a higher ultimate FS (2.0) is applied. The highest steady state acceleration is assumed to be 4.55-g in longitudinal direction and 0.25-g in lateral direction. The internal pressure of the manned AV and the SHM is 101.3 kPa (1 atm). The ECLS section assumes half a bar with 40/60 O2/N2.


Figure 6-50 shows the surface configuration of the assembled LEV:



page 142 of 317

s

Figure 6-50: Lunar Exploration Vehicle

The DM is the unmanned portion of the LEV. To fit in the fairing and to comply with the required footpad distance, a deployable landing system is applied. The stowed DM is shown in Figure 6-51. It consists of four independent legs. The landing load attenuation capability is provided by honeycomb shock absorbers located in the inner part of the legs. Each main leg is connected to a four-legged truss structure, and has lateral stabiliser arms. It is assumed that the maximum vertical landing velocity is 2.5 m/s, and after touchdown within 0.2 s, the velocity will decelerate to 0 m/s. This means that a minimal vertical stroke of about 0.3 m is required for the shock absorbers and that the deceleration will be approximately 13 m/s2. The worst case is an initial touchdown of 1 footpad.

Figure 6-51: Stowed Descent Module

DescentModule

Surface Habitation

Module

LunarAscent Vehicle



page 143 of 317

sThe main structure of the descent stage consists of two pairs of parallel beams arranged in a cruciform, with a platform on the upper and lower surfaces. The outer walls are closed with stiffened side panels. This main structure supports the descent propulsion tank, the SHM and the AV as well. It also provides the attachment for the landing legs and the descent engine. The descent propulsion tank is a load carrying structure and is therefore reinforced by stringers and frames. The tank is connected to the DM main structure by means of a truncated cone and on the upper part it is connected to the SHM by a similar structure.

Figure 6-52: Surface Habitation Module with airlock

For the structure of the SHM (see Figure 6-52) the efficient, lightweight and proven aluminium stringer/skin with frames was selected. A truncated cone provides the structural connection with the AV. A passage tunnel enables the astronauts to move from the AV to the SHM. The AV is shown in Figure 6-53:

Figure 6-53: Ascent Vehicle



page 144 of 317

s6.9.4 Budget

The structure mass budgets of the DM, SHM and AV are shown in Table 6-31, Table 6-32 and Table 6-33, respectively:

Table 6-31: Mass budget of the Descent Module



page 145 of 317

s

Table 6-32: Mass budget of the Surface Habitation Module

Table 6-33: Mass budget of the Ascent Vehicle

[kg] [%] [kg]

SHM_skin 1 738 10 812

SHM_Stiffeners 1 369 10 406

Ring SHM 5 23 10 25

Airlock 1 300 10 330

Interstage Cone to Ascent Vehicle 1 49 10 54

Ring Interstage 2 10 10 11

1591 10 1750

ITEM Nr.M_struct Unit Margin Unit mass with margin

[kg] [%] [kg]

AV_Config_Envelope_3 1 339 10 373

AV_Config_Envelope_3_Stiffening 1 170 10 187

AV_Seat 3 15 10 17

AV_Dock_Tunnel 1 70 10 77

AV_Dock_Tunnel_Stiffening 1 35 10 39

Struts Propellant Tank 8 8 10 9

Platform Propellant Tank 1 57 10 63

Thruster Brackets 4 4 10 4

Small Rings 3 7 10 7

Large Rings 3 18 10 20

869 10 956




page 146 of 317

s6.10 LEV – communications 6.10.1 Requirements and design drivers

The overall LEV data requirements are shown in Table 6-34:

Uplink[Mbps]

Downlink [Mbps]

LEV <-> Earth 17.5 20.2 LEV <-> Earth

during descent/ascent 1.6 6.3

LEV <-> Earth contingencies during

descent/ascent 1.6 0.4

LEV <-> Earth contingencies 10 0.4

EVA <-> LEV 12.6 12.6 Table 6-34: LEV data requirements

For EVA, additionally: • Maximum considered distance to LEV on the lunar surface of 8 km (obtained from

Apollo 17 data) • For simplicity, repeaters shall be avoided and the system shall work also in case of

complex topography, non-line of sight (transmitting and receiving antennas) and multipath environment


6.10.2.1 Band trade-off

With S-band (which is shared by Space Research (SR) Cat. A, Space Operation (SO) and Earth observation services, plus high-density mobile systems), high congestion and sharing difficulties with fixed systems have already been experienced. Therefore S-band will be noisy and the G/S support to it will be reduced or suppressed by ESTRACK in a few years. The most favourable frequency of operation depends on the types of antenna used at both ends of the link (ground and space). Three bands are considered: X (8, 8.4 GHz), Ka (25, 27 GHz) and Ka+ (40, 37 GHz). One of the most straightforward ways of improving link capacity is moving to higher frequency bands. The 40 GHz up and 37 GHz down band (Ka+/Ka+) was allocated by ITU for the very purpose of human space exploration. Contrary to the 34 GHz up / 32 GHz down (Ka/Ka), the Ka+ band can be used for both deep-space (Mars) and near-Earth (Moon) missions, while Ka-band can not be used from the Moon, but from Mars. To use the same frequencies for all human missions, Ka+ band is the best option (see RD[21]). The problems with this band are that Ka+ band is new to space activities and no technology development has been performed so far and secondly, atmospheric and rain attenuation is higher than for Ka-band (approximately 3 dB), so the improvement of gain due to the higher frequency is cancelled by the higher attenuation



page 147 of 317

s Comparing Ka+ and X-band, the following factors are important:

• Assuming constant apertures at both ends, the communication performance can be improved by a factor of 13 dB (theoretical) if the frequency of operations is increased from X to Ka+ band

• The weather dependence of Ka+ band is high, so the availability of the link is lower than X-bands

• Higher pointing accuracy is required, compared with X-band. For example, 4 times more with respect to X-band

• There is more bandwidth availability with respect to X-band (which has only 10 MHz) As an output of the trade-off, both frequencies X and Ka+ will be used: Ka+ will be used as baseline and X as back up because it provides higher availability despite the reduced data rate.

6.10.2.2 High-gain antennas

To avoid the change of orientation of the spacecraft depending on the Earth position, all high-gain antennas will have a steering mechanism, defined by degrees of freedom and pointing precision. The degrees of freedom will be always 2, so any dependence from spacecraft is avoided. As regards pointing precision the worst case assumed is half of beam-width is taken, so a pointing loss smaller than 3 dB will be guaranteed. Ranging is used in the TC and TM signal. The maximum considered distance to calculate the data rate is the worst case, the distance Earth-Moon (400 000 km). A dish diameter of 0.5 m is consistent with the data requirements for Hub and LEV.

6.10.2.3 LGAs

To comply with the requirement of communications in any S/C attitude, three switched LGAs are placed, one for each axis. Isotropic (all directions) coverage will be achieved.

6.10.2.4 X-band communications system

A limitation of 10 MHz for X-band downlink is set in RD[6] for SR 8.4-8.45 MHz and 7.190-7.235 MHz. For a SR Category A mission with a data rate higher than 2 Msps, GMSK BTb=0.25 and Filtered OQPSK are recommended in RD[6]. GMSK BTb=0.25 has a spectral efficiency of 1.16 bits/Hz, while SRCC (α=0.5) OQPSK has 1.5 bits/Hz (see RD[22]). To maximise data transmission, SRCC (α=0.5) OQPSK has been selected due to its higher modulation spectral efficiency. Maximum transmitted symbol rate is 15 Mbps. The received signal-to-noise ratio at the G/S is too low, therefore some coding is necessary. Coding may be selected among the following:

• RS(255, 223) coding, it adds an overhead of 14%, so the final information bit rate is 13 Mbps. This code has a coding gain of 5.8 dB for FER=10-5 and Interleaving 5 (required Eb/N0=6.7dB).

• Convolutional codes ½, 4/3: it adds an overhead of 100%, so the final information bit rate is 7.5 Mbps.



page 148 of 317

s• Concatenated codes: convolutional ½ with RS(255, 223). They add an overhead of

128%, so the final information bit rate is 6.5 Mbps. • Turbocodes ½, ¼, 1/6: they add an overhead of 100%, 300% and 500%. The information

bit rate will be 6.5, 3.25 and 2.5 Mbps, respectively. In conclusion, the coding that minimises the overhead and gives enough coding gain for the link budget is RS(255, 223).

6.10.2.5 Descent/Ascent phases

During descent/ascen the, LEV HGA must be pointing to the Earth G/S to have continuous communications. Due to the LEV movement during these critical phases, pointing accuracy will be the most important factor. The HGA is 0.5 m diameter, and its beamwidth will depend on the band:

• X-band 3 dB beamwidth: 5 degrees • Ka+ band 3 dB beamwidth: 1 degree

The required pointing precision for Ka+ band is approximately 5 times higher than for X-band and data requirements are lower during these phases, so X-band would be enough for downlink data during descent/ascent phases.

6.10.2.6 G/S assumptions

Transmission Reception

Frequency band EIRP Frequency band

Effective G/T, 10º

7145 – 7190 MHz 89.31 dBW (1995W RF) 8400 - 8450 MHz

42.52 dB/K 12-m antenna

Kourou, Perth, Vilspa, Canberra 40-40.5 GHz 84.1 dBW (40 W RF) 37.5-38

GHz 53.06 dB/K

Table 6-35: Assumed ground station characteristics

6.10.2.7 EVA communications

The used transmission protocol/modulation shall be able to match data rate and propagation requirements (resistant to multipath effect and good behaviour with no line of sight communications). Several protocols have been considered (see Table 6-36). Spread spectrum techniques will be the most suitable for obtaining a good behaviour with multipath propagation and no line of sight communications. Among those OFDM will have the best performance and have been selected, in particular the 802.11 g protocol that is commercial (known as WiFi).



page 149 of 317

s

Range

Max Data Rate (raw)

Processing power Multipath resistance Mass Power

Ad-Hoc (direct comms

between computers)

Bluetooth Short (around 10 m)

Medium (1 Mbps,

future 2, 3 Mbps)

High Frequency hopping SS Low Low Yes

802.15.4 Short Low (250 kbps)

Low DSSS Low Low Yes

802.11b (WiFi) Med Med

11 Mbps High FHSS or DSSS Med Med Yes

802.11a/g (WiFi) Med -Long High

(54 Mbps) High 52 sub-carrier OFDM Med Med Yes

UWB * 802.15.3a Long High

200 Mbps Medium Shaped pulse or

Frequency switched OFDM

Low Low Yes?

802.16e* (WMANs) Long High

(75 Mbps) OFDM High? High? No

802.20 * Long Medium (1 Mbps) OFDM Med Med Yes

GSM, CDMA Long Low 9.6 kbps Low Low GSM

Yes CDMA (DSSS) Low Low No

DVB-T/DVB-RCS Long High (19.6

Mbps) OFDM High? Med? No

CCSDS Proximity-1 Long Medium (2 Mbps) Very Low Low Med No

Under development. Expected to be ready in 2-3 years Under development. Expected to be ready in several years

Table 6-36: Possibilities considered for EVA lunar surface communications


6.10.3.1 Communications availability using G/S

During the LEO phase, TDRSS will be used. The link will be established using the Ka-band (22.55 – 23.55 GHz, 25.25 – 27.5 GHz) subsystem in case of LEV or Hub, and S-band in case of PMs. The coverage at the Moon when there is no Moon occultation of the Earth (true only for the Moon Base and for most of the cruise phase), is 22.5 h/day, therefore 93.75% availability. The considered G/S locations are Madrid, Kourou and Perth. Total availability, without considering weather effects, would be achieved by adding a Perth-like G/S in Canberra (Australia).

6.10.3.2 EVA

The main features of the EVA communication system are: • Commercial protocol: 802.11g • No repeaters necessary if there is reflected wave, due to usage of OFDM • Raw data rate: 54 Mbps at 8 km -> approx. 24 Mbps net per channel • All users are interconnected. Maximum of 13 dedicated channels that can be shared • Transmitted power: 1W • Power consumption: 3W • Transmitter/receiver OTS hardware available:

o Mass: 250 g



page 150 of 317

so Size: 10 x 10 x 2 cm

• Omnidirectional systems (in lunar surface plane) with gain 5.2 dBi (see RD[23]) • Processing power needed: EVA suit portable computer needed

See section 4.12 for EVA communications architecture.

6.10.3.3 LGAs data rate

Figure 6-54 shows the uplink/downlink data rate for the X-band link as a function of the range. LGAs will be used in case of contingency with HGAs, if S/C attitude and/or HGA pointing are lost.

Figure 6-54: Uplink / Downlink data rate vs. range for LGA X-band link

6.10.3.4 Descent/ascent phases

Assuming the 0.5-m HGA in X-band or, in case of HGA pointing difficulties, 2 omni-directional LGAs, the link margin is 11 dB, therefore 8 dB can be assigned to pointing precision and 3 dB to the minimum ESA link margin. For 8 dB pointing losses, a pointing precision of ±4 degrees is required (see Figure 6-55). Roundtrip delay (Earth-Moon) is 2.7 seconds, which is low enough to allow the ground segment to participate actively in landing/taking off and docking/undocking operations. For short distances, less than 585 km, UHF can be used for communications between the Ascent/Descent module and Hub for docking/undocking operations (see Table 6-37). For contingencies, a X-band LGA system is provided.



page 151 of 317

s

D Gmax-3dBBEAM3dB

alpha

Pointing Loss vs Pointing Error Angle 0.0 -0.1-0.3-0.5-0.8-1.1

-1.5-1.9

-2.4-3.0

-3.6-4.3

-5.1-5.9

-6.8

-7.7

-8.7

-0.03

-10

-9

-8

-7

-6

-5

-4

-3

-2

-1

0

0.0 0.3 0.5 0.8 1.0 1.3 1.5 1.8 2.0 2.3 2.5 2.8 3.0 3.3 3.5 3.8 4.0 4.3 4.5

Alpha (deg)

Delta

G (d

B)

Figure 6-55: Radiation diagram of 0.5-m X-band HGA

6.10.3.5 UHF link

LEV and Hub will have a symmetrical UHF subsystem. As shown in Figure 6-57 and link details are shown in Table 6-37. For short distances, less than 650 km (for a data rate of 256 kbps), UHF can be used for communications between Ascent/Descent module and Hub for docking/undocking operations. Data rate will depend on distance, as shown in Figure 6-56 (a minimum 3 dB margin shall be maintained). Additionally, it can be used for contingency communications between LEV/AV on the lunar surface and Hub, when a link using the Earth (or relay satellite depending on the landing site) in Ka+ or X-band cannot be used.

LEV/LAV-Hub UHF Link - Range vs Data rate (Kbps)5W RF, -6dBi gain, cancatenated coding

0.001.002.003.004.005.006.007.008.009.00

10.0011.0012.0013.0014.0015.00

0

100

200

300

400

500

600

700

Distance km

Nom

inal

Mar

gin

(dB)

32.00 64.00 128.00 256.00 512.00 1024 2048

Figure 6-56: Data rate compared to distance Hub-LEV/AV



page 152 of 317

s

Figure 6-57: LEV-Hub UHF communications system

6.10.3.6 Communications summary depending on the phase

• LEO: link in Ka band for manned S/C using TDRSS. In case of non-manned S/C, S-band. Inter-spacecraft communications are done using UHF link.

• Cruise: main Ka+ link with G/Ss is used. In case of contingency X-band system is used, with HGA or LGAs depending on the kind of contingency (Ka+ link not working or loss of attitude).

• Moon: similarly to cruise case. However, when on the far side of the Moon, relay satellites will be used. UHF link is used for docking, undocking, ascent and descent manoeuvres for communications Hub-LEV/AV in case of contingency.

• Lunar surface: depending on the landing site, communications using relay satellites in Ka+ or direct link to Earth with Ka+ or X-band (in case of contingency) will be used. UHF with Hub when there is direct visibility

• Moon Descent/Ascent: X-band HGA with Earth and UHF with Hub when there is direct visibility. In case of contingency, omnidirectional LGAs in X-band are used.

Link Ka+ band

1 X-band (HGA) UHF

2 X-band (LGA)

Uplink Downlink Uplink Downlink Uplink Downlink Uplink Downlink

Frequency 40-40.5 GHz 37.5-38 GHz 7.19-7.235 GHz 8.45-8.5 GHz 410-420 MHz 410-420 MHz 7.19-7.235 GHz

8.45-8.5 GHz

Tx power 40W 10W 71W 10W 5W 5W 71W 10W

Modulation NRZ/PSK/PM GMSK. BTb=0.25

SRRC (α: 0.5) Filtered

OQPSK3

SRRC (α: 0.5) Filtered OQPSK

BPSK/PM BPSK/PM SRRC (α: 0.5)

Filtered OQPSK


OQPSK

Coding

Concatenated:Convolutional

+ RS (255, 223)

Turbo Coding ¼ RS (255, 223) RS (255, 223)


+ RS (255, 223)

Concatenated: Convolutional + RS (255, 223)

RS (255, 223) RS (255, 223)

FER Negligible Negligible Negligible Negligible BER=10-5 BER=10-5 Negligible BER=10-6 Bit rate

(worst case) 40 Mbps 100 Mbps 10 Mbps 10 Mbps 256 kbps 256 kbps 10 Mbps 400 kbps

Table 6-37: Links description

6.10.3.7 Antennas

LEV antenna locations are shown in Figure 6-58. All are placed in the AV so they can be used in all phases.

1 Atmospheric attenuation of 4.5 dB, this corresponds with an elevation of 10 degrees and G/S availability of 90%. 2 Max distance Hub-LEV/AV of 368 km, corresponding to a Hub elevation over the horizon of 10 degrees. 3No standard modulation exists for uplink high data rates.



page 153 of 317

s

Figure 6-58: Antenna locations


6.10.4.1 Antenna summary

Table 6-38: Antenna summary

6.10.4.2 Equipment and mass summary

Element 3 Unit Namecell name

Click on button below to insert new unit

1 e3_unit1_name UHF omni antenna 3 1.2 Fully developed 5 3.82 e3_unit2_name Ka+/X-band dish antenna 2 2.5 Fully developed 5 5.33 e3_unit3_name UHF transceiver 3 0.8 To be modified 10 2.64 e3_unit4_name Ka+/X-band transponder 3 4.0 To be developed 20 14.45 e3_unit5_name SSPA Ka+ band 3 0.8 To be modified 10 2.66 e3_unit6_name SSPA X-band 3 0.8 Fully developed 5 2.57 e3_unit7_name Global RFDU unit 1 10.0 To be developed 20 12.08 e3_unit8_name X-band omni antenna 2 0.4 Fully developed 5 0.89 e3_unit9_name UHF diplexer 1 0.4 Fully developed 5 0.410 e3_unit10_name Fully developed 5 0.011 e3_unit11_name Fully developed 5 0.012 e3_unit12_name To be modified 10 0.0- Do not use To be developed 20 0.0

9 39.0 14.1 44.5

Margin Total Mass incl. margin

Element 3: Lunar Ascent Vehicle MASS [kg]Unit Quantity Mass per

quantity excl. margin

Maturity Level

Click on button below to insert new unitELEMENT 3 SUBSYSTEM TOTAL

Table 6-39: AV equipment summary

Element 2 Unit NameClick on button below to insert

new unit1 EVA 802.11g router/rxer 2 0.3 To be developed 20 0.72 EVA 802.11 antenna 2 0.1 To be developed 20 0.2- To be developed 20 0.0

2 0.8 20.0 1.0

MASS [kg]Mass per


Maturity Level Margin Total Mass incl. margin

Click on button below to insert new unitELEMENT 2 SUBSYSTEM TOTAL

Element 2: Surface Habitation ModuleUnit Quantity

Table 6-40: SHM equipment summary

Element Type of antenna Band Pointing device Transmitted

power (W)

LEV/AV

Two dish 50 cm Two LGA antennas (quadrifilar helix) Three omnidirectional UHF antenna

Ka+ and X-band X-band UHF

2 DOF Steering None None

10 10 5



page 154 of 317

s6.10.5 Options

6.10.5.1 EVA communications: UWB

The option for OFDM is UWB, based on a completely different transmission method. UWB emits rapid sequencing of extremely short (<1 ns) wideband (>1 GHz) low-power bursts of radio frequency energy. UWB will reduce power, mass and volume over conventional communications systems. Maximum data rates are expected to be around 200 Mbps for close distances.

6.10.5.2 Experimental optical return link

In the present scenario of Martian exploration within the Aurora programme, an optical system is foreseen to achieve high-link capacity in the Mars-Earth link. Though not strictly required, a lunar mission would be an opportunity to validate such technology.

6.11 LEV – data handling 6.11.1 Requirements and design drivers

6.11.1.1 Functional requirements

The LEV data handling system needs to be capable of performing the following functions with a high level of autonomy:

• AOCS/GNC. Interface the sensor equipment and control the propulsive system during all mission phases. Perform calculations of trajectories and manoeuvres.

• Telecommands. Include a TCs handler that demodulates, decodes, validates, distributes and executes both real-time and time-tagged ground commands, with the supervision and control of the crew on board.

• Telemetry. Acquire housekeeping data for transmission to ground and internal processing and support functions for payload data management: Acquisition of payload data and support basic data processing, data storage and data transmission to ground.

• Power control. Monitor the battery status (mainly the charge/discharge current and voltage) and the solar arrays.

• Thermal control. Keep the vehicle temperature inside definite limits by reading thermal sensors and control heaters.

• Life support control. Monitor and control the life-support variables levels (oxygen, temperature, etc) inside the spacecraft.

• On-board time. Provide on-board time reference generation and distribution to ensure synchronisation and time tagging of attitude data for post-processing.

• Failure detection and recovery. Include functions to reconfigure faulty elements and functions for restoring the vehicle to a nominal state. Contingency management.

• On-board storage. Provide on-board data storage capability to store all housekeeping and payload data.

• Human-machine interfaces. Provide interfaces on-board for the crew to monitor and control the spacecraft.

• EVA computers. Provide portable computer equipment to be integrated in the spacesuit.



page 155 of 317

s6.11.1.2 Performance requirements

To be able to perform all the above-mentioned functions, several performance aspects need to be analysed. The DHS processing units required capabilities include: having to perform real-time data processing, involving complex algorithms for guidance, navigation and control, while keeping operational all the other functions. The system shall allow the crew to be informed of the current situation at every moment. The computers and human-machine interfaces need to be so that a very short reaction time can be achieved, in case the crew would need to manually take control of the situation.

6.11.1.3 Safety requirements

All the required functions need to be achieved while being able to guarantee the maximum possible level of safety for the crew. From the DHS point of view, this safety implies two main design drivers: reliability and availability.

6.11.2 Main issues and proposed building blocks

There are several aspects to consider when assessing the required building blocks for the data handling system of such a mission. The main issues concern computers and interconnectivity performance as well as safety, reliability and availability. There are other issues that are a direct consequence of the specific characteristics of manned spaceflight: increased software complexity, provisioning and maintenance of components and units to support a long-duration programme, development of appropriately advanced human-machine interfaces and EVA computers, etc. Some of these main issues are hereafter analysed in more detail and some building blocks have been identified that would be required to achieve a sustainable lunar exploration programme.

6.11.2.1 Computing power

The present and near-future European space computers are based on the SPARC RISC architecture (ERC32, LEON2FT, and soon LEON-3). More powerful rad-hard Power-PC are currently available, but under US ITAR control. The computing power of current and future ESA microprocessors is shown in Figure 6-59. LEON-2 is capable of 100 Mips and will be available in 2006. The first flight models of LEON-3 are expected to be available in 2009 and will provide 200 Mips. The LEON-3 design will be optimised for implementing a multi-processor architecture, up to four LEON-3s on a single chip. Theoretically this would allow for an enhanced performance of up to 800 Mips. However, to exploit this capability, the software development process and tools need to be adapted accordingly. Note that US processors available today already provide 240 to 300 Mips, and a 370 Mips version is foreseen before 2007. It is assumed that the processing power capabilities of the LEON-2 processor would be enough to cope with the basic mission requirements. However, considering the current roadmap, LEON-



page 156 of 317

s3-based computers will be available enough before the first launches of the lunar exploration programme so this is selected as baseline. The use of COTS solutions (mainly Power PC) shall be assessed and is proposed as an option for non-critical payloads. The availability of SOI process may boost safe use of high-power chips in space, but dedicated development is needed.

Figure 6-59: ESA roadmap for microprocessors

6.11.2.2 On-board interfaces (communications and interconnectivity)

Standardisation of interfaces started in commercial electronics mainly for peripheral interconnection. Now through ISO and IEEE, it is involving practically any part of the design. The current avionics are based on well-established standards as regards ground-spacecraft communications (CCSDS, ECSS) and hardware-level discrete interfaces (ESA TTC-01B). The future standards that will cover all the interfaces and protocol layers required on board, to ensure the independence between hardware and applications, are still in development. The avionics architecture should be an open architecture based on well-known industrial standards acting at all layers, from hardware to embedded software to human interfaces. The internal communications and interconnections for the different levels of the design are especially important to ease the subcontracting of avionics subelements whilst minimising the risk at integration level. Among other standardisation efforts, the CCSDS SOIS (Spacecraft On-Board Interfaces) (see RD[45]), providing isolation between the S/W applications and the communications logical and



page 157 of 317

sphysical layers can be seen as the base common language towards a distributed intelligence spacecraft, capable of having functional redundancies so as to increase the system reliability. The outcome of standardisation of on-board interfaces will be a set of buses covering all the levels of modular hardware integration. Each interconnection type is specialised and a stack of interconnection means, hierarchically organised, is necessary to guarantee functional redundancy and system optimisation. The SOIS layers will be in charge of hiding the system hardware complexity at application software level.

Figure 6-60: Peripheral interconnection buses versus speed and complexity of the terminals

Figure 6-60 shows several present interconnection buses, some of them already qualified (shown in green in the figure) for space. Other widely used terrestrial technologies will soon be available for space use, such as the very popular wireless communications. As a particular enabling technology for this type of mission, it will facilitate the interaction of the crew with the control systems, and will enable high-quality data, video and voice communications during Extra Vehicular Activities (EVA) on the lunar surface. Other important contributors for a more structured data network on board are sensor buses such as OneWire or I2C. These low complexity and low-speed buses will permit the collection of housekeeping data such as thermal control, with digital multi-drop lines, drastically reducing the harness mass usually required in present missions for hundreds of point-to-point lines.

6.11.2.3 Safety: reliability and availability

A high reliability is generally achieved avoiding “single-point failures” by increasing the number of redundant units. However, duplicating or triplicating units as such would have a considerable impact on the mass budget. The architecture proposed is therefore based on functional redundancy. There will be several processing units available in the spacecraft, and in principle, each one will have some specific



page 158 of 317

sfunctions allocated. Not all of these functions will be required at all times, so in case of a failure occurring in one of the units, any other processing unit can be reconfigured to perform the required functions. To provide high availability, every processing unit is internally triple redundant with majority voting at the outputs. This will be required for critical phases, such as docking or landing, when delay for computer reconfiguration is unacceptable if an error occurs. The network is also a very important element of the architecture, and therefore, single-point failures need to be avoided. The approach for network redundancy is different for low- and high- speed networks as shown in Figure 6-61. The CAN bus architecture has four redundant buses, and the approach is the one used in ATV. For the high-speed network, redundant routers are suggested, as well as an architecture where alternative network paths would be available, to be able to cope even with double failures.

Figure 6-61: Redundancy architectures for CAN bus and SpaceWire networks

An independent network for caution & warning data distribution has to be considered to generate and communicate the appropriate alarms. It is also important to define which power lines need to be independent to avoid single-point failures in the power lines. Another important safety requirement is that non-inhabited modules shall be controllable remotely from inhabited modules. The appropriate interconnection has to be established then between the different modules when docking.

6.11.2.4 Software complexity

The requirement for functional redundancy and also for triple modular redundancy with majority voting architectures has an important influence on software design. Increased software modularity and portability is needed to provide the required flexibility. Also a high level of autonomy is demanded for this kind of mission, involving enhanced capabilities, such as on-board calculation of trajectories and manoeuvres, contingency management, FDIR, etc.



page 159 of 317

sThe system architecture is designed to enable the decentralisation of functions and the portability and exchangeability of functions among the different processing units. While not yet being a completely distributed architecture, with transparent distribution of functions, the architecture proposed already implies a non-negligible software complexity. A set of well-established standards and protocols becomes really useful when providing standard software building blocks and to hide all the complexity of the underlying software and hardware layers at application level. CCSDS File Delivery Protocol (CFDP) is a good example of a new international standard, developed by the Consultative Committee for Space Data Systems (CCSDS) to meet a comprehensive set of deep-space file transfer requirements as articulated by a number of space agencies including NASA, ESA, NASDA, CNES and BNSC/DERA. In addition, CFDP will serve as a prototype for the future Interplanetary Internet as envisioned by the IPN study team (see RD[47]).

6.11.2.5 Maintenance/obsolescence

The availability of space-qualified electronic components has decreased in the last years. Moreover, the evolution of rad-hard technologies after 0.13 µm is very difficult to predict. It is not yet known whether the hardening techniques at design level will permit compensating the increasing SEE sensitivity. Consequently, the gap between commercial technologies and rad-hard technologies may decrease or increase. This introduces a problem in performance. Already in some present designs, sometimes basic EEE parts are chosen not for architecture optimisation but because the lack of any other choice. Currently (even including ITAR licenses), there is a severe lack of choice in ADCs, analogues, power components, FPGAs and memories. To support a long-duration programme, provisioning of components and units is important to avoid having obsolescence problems. It might be necessary to provide dedicated electronic supply lines to support such an effort. The evolution towards more modern production lines must also be encouraged supporting European foundries (Atmel and ST) up to the newest SOI processes.

6.11.2.6 User interfaces - EVA equipment

A very important aspect is the requirement for appropriate human-machine interfaces for the crew to be able to monitor and control the spacecraft. They should also provide interfaces for the payload equipment, as the crew needs to interact with them to achieve the scientific objectives of the mission. It is also a safety requirement that controls should be operable by a pressure-suited astronaut, and this needs to be taken into account for the design of appropriate human-machine interfaces, especially for critical controls.



page 160 of 317

s

Figure 6-62: Wearable computers

In particular for Extra Vehicular Activities, the crew must be supported by adequate computer equipment. There is no such technology already available for space, but some developments for terrestrial applications (such as described in RD[48]), provide very interesting features for space such as the wearable computers shown in Figure 6-62. These small computers with low power, low mass (500 g), and reduced dimensions 150x90x50 mm, provide interesting human-machine interfaces such as voice activation, touch screen, wrist-wearable keyboards that shall be considered for this kind of mission. In particular, a voice-commanded computer seems a good solution, other interfaces need to be adapted to be usable by a pressure-suited astronaut in a dusty environment. High-speed wireless interfaces will be required to transmit data, video and voice to the surface habitation module while in EVAs. Some commercial standards have been identified to establish the first lunar wireless hot-spot.


A fully decentralised data handling system has been proposed. Resources must be distributed but well located. The network then becomes the key DHS component. Network design becomes a critical aspect, it being very important to avoid undesired bottlenecks and to keep control of data flow. The network is composed of several levels, which ensures the best possible performance for each type of traffic. The main backbone network proposed has two levels: a low-speed command and control network, and a high-speed data network. Following a hierarchical approach, other lower- speed wired and wireless sensor networks are expected to achieve a low harness mass. High- speed wireless networks are also proposed for internal use to facilitate the interface to the user and for external use to increase mobility in lunar surface operations. To allow for functional redundancy, several computers must be able to perform any task, so they have to present similar characteristics. They can be reprogrammed from the safeguard data recorder. The housekeeping units are independent and accessed through the network. The TM/TC interfaces are also independent and accessed through the network. In this way, any computer unit duly reprogrammed can access the housekeeping data from the satellite, receive telecommands and send telemetry to ground, acting as the main control unit. Considering the LEV subdivision into three modules: Descent Vehicle, Surface Habitation Module, and Ascent Vehicle, the architecture has been optimised to have the minimal equipment



page 161 of 317

spossible in the ascent vehicle. In total, three data processing units are proposed, each of them internally triple redundant with majority voting, with the following distribution of functions:

• Descent/Ascent control unit. This unit is in charge of controlling the spacecraft descent and ascent AOCS/GNC control loops. It is therefore situated in the Ascent Vehicle, but is also used during descent phase.

• GN&C computer. This unit is in charge of controlling the spacecraft GNC, during cruise, orbital and other phases than descent/ascent. It also acts as a spare unit that could replace the Descent/Ascent control unit in case of failure of this one. It is also located in the Ascent Vehicle.

• General System Computer. This unit is in charge of performing the necessary payload data processing during all the mission phases until surface operations. It is situated in the Descent Module and will be left on the lunar surface. It can also replace any of the other mentioned computers. As an option, and if a higher processing power is required to handle payload data, this unit could be based on COTS processors, using TMR/EDAC techniques for increased reliability. In this case, it would be recommended to have another spare unit with similar characteristics as the two above mentioned units, to serve as functional replacement in case of failure.

Two different types of memory units have been considered in the design:

1. The Safeguard Data Recorder will be in charge of keeping the software modules that need to be executed in the different processing units. It may also be used for some critical data

2. The Mass Memory Unit proposed is already quite a standard development today. Future technology will increase the ratio of memory capacity / mass & power. It will be used to temporarily store housekeeping, payload and scientific data prior to sending to ground. It will allow storing up to 8 hours of data during a hypothetical communications outage, at an average data produced of 6 Mbps

The backbone network interconnection (both low and high speed) of the different modules has to be ensured at the moment of docking or immediately after. Figure 6-63 and Figure 6-64 show the data handling system architecture. The Lunar Descent and Surface habitation modules have been considered together. In the figure, the wireless access point for the surface operations is shown, to show an example of the connectivity for the EVA computers and other optional handheld devices.



page 162 of 317

sSurface Habitation Module

Micro RTUs UNITS 10.00dim 133.00 95.00 136.50 mm3mass 4.00 kgpow 8.00 W MMI+SCREEN

Computerkeyboard

Wireless Access pointdim - - - mm3 SPACESUITmass - kg Computerpow - W

Wireless I/F

SPACESUITComputer

Wireless I/F

HANDHELDDevice

Wireless I/F

Lunar Descent Module

Descent/Ascent control unit Safeguard Data Recorderdim 266.00 190.00 210.00 mm3 dim 266.00 190.00 126.00 mm3mass 8.30 kg mass 4.90 kgpow 35.00 W pow 5.00 W

DM on board bus - 1 mbit

General System Computer Mass Memory Unitdim 266.00 190.00 273.00 mm3 dim 266.00 190.00 126.00 mm3mass 10.85 kg mass 4.90 kgpow 29.89 W pow 18.33 W

Analog I/OHousekeeping Unitdim 266.00 190.00 210.00 mm3 Digital I/Omass 8.30 kgpow 6.02 W Serial Lines

Communication Unitdim Included in General System Comp Transponders --> Communications Subsystemmass But with an independent BUS I/Fpow

Micro RTUs UNITS 10.00 MMI+SCREENdim 133.00 95.00 136.50 mm3 Computermass 4.00 kg keyboardpow 8.00 W

Distributed control busspacecraft backbone MMI+SCREEN

ComputerHigh Speed Data Network keyboardCommand & Control Network

Figure 6-63: DHS architecture of Surface Habitation and lunar Descent Modules



page 163 of 317

sLunar Ascent Vehicle

Descent/Ascent control unit Safeguard Data Recorderdim 266.00 190.00 210.00 mm3 dim 266.00 190.00 126.00 mm3mass 8.30 kg mass 4.90 kgpow 35.00 W pow 5.00 W

DM Internal Bus

GN&C Computer bus towards GNC systemsdim 266.00 190.00 210.00 mm3mass 8.30 kgpow 26.16 W

Communication Unitdim Included in GN&C Computer Box Transponders --> Communications Subsystemmass But with an independent BUS I/Fpow

Analog I/OHousekeeping Unitdim 266.00 190.00 210.00 mm3 Digital I/Omass 8.30 kgpow 6.02 W Serial Lines

MMI+SCREENComputerkeyboard

Micro RTUs UNITS 5.00dim 133.00 95.00 136.50 mm3 MMI+SCREENmass 2.00 kg Computerpow 4.00 W keyboard

Figure 6-64: DHS architecture of Lunar Ascent Vehicle

The computers proposed are based on LEON-3 processors that provides a performance of 200 Mips. The high-speed network is SpaceWire based (300 Mbps) and the low-speed proposed is based on CAN (1 Mbps). The important elements of the architecture are the micro remote terminals units proposed. They help decentralise resources, providing distributed access to the backbone network with an important gain in units and harness mass, power and integration ease, with respect to other more traditional centralised RTUs approaches. This approach is already the baseline for future spacecraft data handling system developments.

6.11.4 Data budgets

The first two tables of Figure 6-65 show the estimated data requirements for the communications with ground in nominal and contingency cases. The third table shows the data requirements for the communications between ground and LEV during the descent and ascent phases. The last one shows the requirements for communications between the LEV module and the EVA Computer systems.



page 164 of 317

s Data Communications - Earth <-> LEV/Hub

Total Trans same time Rate Kbps U Rate Kbps D Total rate U Total rate D Duty Cycle Sus BERHQ Video transmission channels 10 3 4000.00 4000.00 12000 12000 5% 1.E-04LQ Video Trans Channels 50 10 128.00 128.00 1280 1280 5% 1.E-05Audio transmission channels 50 10 64.00 64.00 640 640 25% 1.E-06Earth Data Network Links 40 10 256.00 512.00 2560 5120 100% 1.E-09System Housekeepings 1000 300 0.51 0.51 154 154 100% 1.E-08

Uplink DownlinkMaximum Data Rate (overall) 16634 19194Average Data Rate (overall) 3538 6098Minimum Data Rate (overall) 154 154

Data Communications - Earth <-> LEV/Hub ContingencyTotal Trans same time Rate Kbps U Rate Kbps D Total rate U Total rate D Duty Cycle Sus BER

HQ Video transmission channels 10 1 4000.00 4000 0 5% 1.E-04LQ Video Trans Channels 50 10 128.00 1280 0 5% 1.E-05Audio transmission channels 50 2 64.00 64.00 128 128 25% 1.E-06Earth Data Network Links 40 1 2048.00 64.00 2048 64 100% 1.E-09System Housekeepings 1000 300 0.51 0.51 154 154 100% 1.E-08

Uplink DownlinkMaximum Data Rate (overall) 7610 346Average Data Rate (overall) 2498 250Minimum Data Rate (overall) 154 154 Data Communications - Earth <-> LEV Descent/Ascent

Total Trans same time Rate Kbps U Rate Kbps D Total rate U Total rate D Duty Cycle Sus BERHQ Video transmission channels 10 1 4000.00 0 4000 5% 1.E-04LQ Video Trans Channels 50 3 128.00 128.00 384 384 5% 1.E-05Audio transmission channels 50 6 64.00 64.00 384 384 25% 1.E-06Earth Data Network Links 40 2 256.00 512.00 512 1024 100% 1.E-09System Housekeepings 1000 300 0.51 0.51 154 154 100% 1.E-08

Uplink DownlinkMaximum Data Rate (overall) 1434 5946Average Data Rate (overall) 781 1493Minimum Data Rate (overall) 154 154 Data Communications - LEV <-> EVA (Wireless 802.11g)

Total Trans same time Rate Kbps U Rate Kbps D Total rate U Total rate D Duty Cycle Sus BERHQ Video transmission channels 1 1 4000.00 4000.00 4000 4000 5% 1.E-04LQ Video Trans Channels 2 2 128.00 128.00 256 256 5% 1.E-05Audio transmission channels 5 1 128.00 128.00 128 128 25% 1.E-06Data Network Links 30 15 512.00 512.00 7680 7680 100% 1.E-09Biomedical/Space Suit Data 1000 300 0.51 0.51 153 153 100% 1.E-08

Uplink DownlinkMaximum Data Rate (overall) 12217 12217Average Data Rate (overall) 8078 8078Minimum Data Rate (overall) 153 153

Figure 6-65: Data communication requirements



page 165 of 317


Element 3 Unit NameClick on button below to insert new

unit1 Descent/Ascent control unit 1 8.3 To be developed 20 10.02 GN&C Computer 1 8.3 To be developed 20 10.03 Safeguard Data Recorder 1 4.9 To be developed 20 5.94 Housekeeping Unit 1 8.3 To be developed 20 10.05 Micro RTUs 1 2.0 To be developed 20 2.46 Man Machine Interface 2 5.3 To be developed 20 12.7- To be developed 20 0.0

6 42.4 20.0 50.8ELEMENT 3 SUBSYSTEM TOTAL Click on button below to insert new unit




Figure 6-66: Ascent Vehicle list of equipment


unit1 General System Computer 1 10.9 To be developed 20 13.02 Safeguard Data Recorder 1 4.9 To be developed 20 5.93 Mass Memory Unit 1 4.9 To be developed 20 5.94 Man Machine Interface 3 5.3 To be developed 20 19.05 Housekeeping Unit 1 8.3 To be developed 20 10.06 Cameras 4HQ + 20LQ 1 1.8 To be developed 20 2.27 Micro RTUs 1 4.0 To be developed 20 4.88 EVA Computers 4 1.0 To be developed 20 4.8- To be developed 20 0.0

8 54.6 20.0 65.5ELEMENT 2 SUBSYSTEM TOTAL




Mass per quantity excl.

margin

MASS [kg]

Figure 6-67: Surface Habitation Module and Descent Vehicle list of equipment

6.12 LEV – mechanisms 6.12.1 Requirements and design drivers

As a result of the LEV’s configuration, the following necessary mechanisms and their requirements can de derived:



page 166 of 317

s

Vehicle Mechanism Requirement

Descent Module Deployment and Locking system for the Deployable

Legs

X 4 legs

Locking Latches for the Landing Leg in the stowed

configuration

X 4 legs

Surface Habitation Module Crew Egress Hatches Hatch diameter 0.9 m

Vehicle Stage Separation System

I/F diameter 2 m

EVA Egress Hatches - Internal and External

Hatch diameter 1.2 m

Inter-stage berthing for LDM-SHM I/F

Solar Panel Hold-down & Deployment

Panel area 15 m2

Lunar Ascent Vehicle External Hatch- AV-Hub & AV-SHM

Hatch diameter 0.9 m

Berthing & Docking in Lunar Orbit- AV to Hub

Un-docking during Lunar Orbit- AV to Hub

Antenna Pointing and Tracking Mechanism Surface

to Aero-stationary satellite communication antenna

Antenna diameter: 0.4 m

Antenna mass: 2.5 kg

Coverage:180° hemispherical.

Pointing accuracy: 10°

Boom mounted at 0.5 m

Number of antennas: 2

Table 6-41: LEV mechanism requirements



page 167 of 317

s6.12.2 Baseline design

6.12.2.1 Landing System Deployment and Lock (LDM)

This device will block the damping system of the legs due to re-entry loading. No specific design has been performed. A mass estimate is entered based upon the assumption of a spring-actuated system.

6.12.2.2 Crew Egress Hatches- SHM & AV

Sealable hatches are required for the following I/Fs: • Hub to AV I/F through the IBDM • SHM to AV • SHM EVA hatches as part of airlock - internal and external

The hatch size, with the exception of the airlock hatches, is chosen to suit the IBDM available internal diameter (∅=0.9 m). The hatch will require latch and seal mechanisms. Mass estimates have been also entered using a ‘simple geometry’ model and material mass properties. The airlock hatch diameter is sized to ≈∅1.2 m to allow for a fully suited astronaut to pass through the outer hatch. The hatch will require latch and seal mechanisms. Mass estimates have been entered using a “simple geometry” model and material mass properties.

6.12.2.3 Vehicle separation SHM-AV I/F

The separation of the MAV from the SHM prior to launch shall be performed by a pyrotechnic operated clamp-band which is estimated to be ∅2 m.

6.12.2.4 Berthing capability

For in-orbit assembly of the LDM to the SHM/AV composite, the Common Berthing Mechanism (CBM) shall be the baseline adapted to suit the purpose of a non-pressurised inter-vehicle connection.

Figure 6-68: Common Berthing Mechanism



page 168 of 317

s6.12.2.5 Solar panel deployment systems

The deployment of the panels shall be performed by a motor-actuated hinge system (two hinges per panel). The deployment support beam structure will aid the support of the panel in the lunar gravity environment. The required 15 m2 will be split into two deployable wings, each with 2 x 5 m2 (5 m x 1 m) panels. The total 20 m2 provided will have a slightly reduced active area on the inner panel. Per panel, 10 hold-down points are envisaged combined with hard snubber points to control the vibration characteristic of the panel. A typical system could be based upon the thermal knife or TiNi actuator/Frangi bolt for the release function to minimise release shocks.

6.12.2.6 Communication antennas

Current APM systems are able to meet the pointing requirements for the antenna. A suitable unit has been chosen and an estimate of the deployable boom mass has been made. Note, however, that the capability of this unit is better than required and therefore some modification and optimisation of the unit can be expected. A schematic of a typical boom mounted deployment and pointing mechanism for the TV antennas is shown in Figure 6-69:

Deployment Hinge

Axial Hinge

Azimuth Hinge

Antenna Mounting I/F

Figure 6-69: Antenna deployment mechanism

For complete 180° hemispherical coverage, the axial and azimuth axes shall rotate through 180°.

6.12.2.7 AV docking

The International Docking and Berthing Mechanism shall be implemented for the AV to Hub interface.



page 169 of 317

s

Figure 6-70: IBDM

The IBDM in its current form has the following mechanical characteristics based upon design analysis:

• OD 1.371 m; ID 0.813 m (effective pass through diameter); H 0.254 m (retracted) • Mass estimate 304 kg • Interface loads (at the sealing interface)- acting simultaneously while docked (flight-

limit): o Axial load (1200 lbf) 5338 N o Shear load (1000 lbf) 4448 N o Bending moment (80000 in*lbf) 9039 Nm o Torsion moment (70000 in*lbf) 7909 N*m

• Internal pressure (16 psi) 110.3 kPa • Life: 15 years, functional life: 20 berthing/un-berthing or docking/undocking cycles

The advantages of the IBDM system are:

• It is an androgynous system implemented on both sides of the mated interface (no male to female connections)

• Due to the system being present on both sides of the I/F, a fully redundant system is realised

Currently, further IBDM development is focused upon increasing the available internal diameter to enable racks to be passed through the IBDM. The intended internal diameter will increase to between 1.0 and 1.2 m (from the current 0.81 m) without increasing the structural ring dimensions by placing all the structural connections outside the pressurised volume rather than inside. To calculate an estimated mass, the ratio of the internal diameter increase shall be used to estimate the increase in the mass. The changes to the IBDM are not major structural changes, so 50% of the calculated increase shall be added to the original mass to give the mass estimate for the increased size IBDM. The estimated mass for this IBDM type is 380 kg. Since a total of five IDBMs are used in the system and commonality is sought in the system, the increased size IBDM has not been implemented due to the resulting mass penalty/increase. The IBDM function is supported by eight electronic boxes of approx. size 0.4 m x 0.25 m x 0.25 m with a mass of 8 kg each.



page 170 of 317

s6.12.2.8 Deployed appendage loads

Due to the lunar gravity environment, the loads on the deployed appendages have been assessed. These are reported in Table 6-42 and Table 6-43:

Solar Array Wing MassDeployment System (Kg) Panel/Cells (Kg) Total (Kg)

16 30 46CoM position 1.25 mAllowable Shear 148 N (Derived from Bending allowable, not actual shear)Allowable Bending 185 Nm

Acceleration (m/s)

SA Force applied at CoG/Shear

LoadSA Bending

MomentMargin of Safety

Bending

Surface Operations Deploying/Deployed 1.62 74.52 93.15 0.986044015(Moon's Gravity Constant)

Table 6-42: LEV solar array loads Communication AntennasHGA-1 HGA-2Mass- Dish 3 Kg Mass 2.5 KgMass APM 9.4 Kg Mass APM 9.4 KgTotal Mass 12.4 Kg Total Mass 11.9 KgDia 0.5 m Dia 0.5Boom Length 0.5 m Boom Length 0.5 Kg

Acceleration (m/s)

HGA-1 Antenna Force applied at

CoG/Shear Load (N)HGA-1 Bending Moment (Nm)



Surface Operations- Deployed 1.62 20.088 10.044 19.278 9.639 Table 6-43: LEV antenna loads

6.12.2.9 Budgets

Element 1 Unit Name


1 Common Berthing Mechanism- Passive 1 315.0 10 346.52 Landing Leg Deployment Mechanism 4 10.0 20 48.03 20 0.0- 0.0 20 0.0


Unit QuantityElement 1: Lunar Descent Module



margin

MASS [kg]


Table 6-44: Descent Module mechanism mass budget

Element 2 Unit NameClick on button below to insert new unit

1 Hatch Door- Airlock 2 38.0 20 91.22 Hatch Door Locking Mechanisms- Airlock 2 170.0 20 408.03 Clamp-band- LAV I/F 1 15.6 20 18.84 Common Berthing Mechanism- Active 1 235.0 20 282.05 Solar Array Panel Holddown systems 4 0.20 10 0.96 SA Deployment Mechanisms 2 13.7 10 30.27 SA SDM/panel Hinhes 2 0.7 10 1.58 SA Yoke Panel 2 4.0 10 8.89 SA Root Hinge 2 5.0 20 12.0

10 SA Drive Mechanism 2 4.5 20 10.811 SA Drive Mechanism Electronics 1 4.5 20 5.412 Hatch Door- Extrenal Egress 1 15.0 20 18.013 Hatch DoorLocking Mechanisms- External Egress 1 75.0 20 90.0- 20 0.0






MASS [kg]

Table 6-45: Surface Habitation Module mechanism mass budget



page 171 of 317

sElement 3 Lunar Ascent Vehicle

Click on button below to insert new unit1 Docking Mechanism- IBDM 1 334.4 10 367.82 Electronic Box- IBDM 6 8.8 10 58.13 Hatch Door- Egress External 2 15.0 20 36.04 Hatch Door Locking Mechanisms- Egress External 2 75.0 20 180.05 Antenna Pointing Mechanism 2 8.5 20 20.46 Antenna Deployment Boom 2 1.0 20 2.47 Antenna Pointing Mechanism Electronics 2 5.0 20 12.0- 20 0.0

7 596.2 13.5 676.7ELEMENT 3 SUBSYSTEM TOTAL Click on button below to insert new unit


quantity excl. Margin Total Mass

incl. margin

Table 6-46: Lunar Ascent Vehicle mechanism mass budget

6.13 LEV – power 6.13.1 Requirements and design drivers

6.13.1.1 Mission requirements

The Power Subsystem shall be sized for 14 surface days (including a contingency duration in case of problems). Moreover, the Lander has to stay ready to take off during the whole stay on the surface. Due to the complexity of a night survival (mainly for the power and the thermal subsystems) design, the surface stay shall only take place during the sunlight: the lunar nights and the Sun eclipses by Earth are avoided. The crew of the LEV is composed of three members for a mission planned to start between 2020-2025 with technologies expected to be ready and qualified in 2015. The descent phase is starting from the separation of the Hub until landing on the lunar surface. The ascent phase is starting from take-off until docking to the Hub. In both phases, the LEV shall be sized for a duration of 24 hours even if the nominal duration is only several hours. Concerning the landing site, the requirement is to make the LEV design able to cope with any latitude. As a starting point, a landing around the equator is assumed; as a second step, the possibility of performing missions at higher latitudes is assessed.

6.13.1.2 Lunar environment description

Figure 6-71 shows the worst-case Sun illumination during the sunlight period of 14 days at the equator and at a 29º latitude location,. During the first and the last days of the sunlight period, the Sun’s elevation is low, while around day 7, the Sun is close to the local zenith.



page 172 of 317

sWorst Case Solar Illumination on the ground

0

200

400

600

800

1000

1200

1400

0 2 4 6 8 10 12 14 16

Days

W/m

2

Equator29deg Worst Case Illumination

Figure 6-71: Worst-case solar illumination on the lunar surface

On the Moon, the sunlight is only diffused by the dust lifted by landing and operations performed by the crew. In some Apollo missions, a strong dust effect following the landing was reported. A dedicated study of the Moon dust properties should be done to limit the dust attenuation as possible and to make an accurate assessment of the resulting degradations. In the frame of this study, possible dust removal mechanisms have been assessed that could be implemented for the LEV.


6.13.2.1 Power budget

The overall LEV power requirements are shown in Table 6-47:



page 173 of 317

s

Table 6-47: LEV power requirements

Including the harness losses and excluding the dissipation of the power subsystem itself, the average bus power level is about 1.5 kW on the lunar surface and 900W during flight mode. In terms of energy required per mode:

• 25 kWh are required for the descent phase • 457 kWh for the surface stay • 18 kWh for the ascent phase

A conservative margin of 20% on top of these values is applied for the sizing of the power units.

6.13.2.2 Technologies trade-offs

6.13.2.2.1 Primary energy sources

The following four candidates for power generation have been identified: 1. Fuel cells 2. Photovoltaic conversion 3. Solar dynamic 4. Thermoelectric generators



page 174 of 317

s6.13.2.2.1.1 Fuel cells

A fuel cell is an electrochemical device that combines hydrogen fuel and oxygen from the air to produce electricity, heat and water.

Figure 6-72: Primary fuel cell with air and hydrogen as reactants

There are five types of fuel cells: 1. Phosphoric Acid Fuel Cell (PAFC): Phosphoric acid is used as an electrolyte. It needs to

operate around 200ºC and the cathode performance is inefficient. PAFCs are usually large and heavy fuel cells.

2. Proton Exchange Membrane Fuel Cell (PEMFC): The mechanism is the same as PAFC but these cells operate at lower temperatures (about 100ºC). They have a high power density and can vary their output quickly to meet shifts in power demand.

3. Molten Carbonate Fuel Cells (MCFC): an alkali metal carbonate (Li, Na, K) is used as the electrolyte. It needs to operate at about 600ºC.

4. Solid Oxide Fuel Cells (SOFC): solid, nonporous metal oxide electrolytes are used. The cell operates at about 800-1000ºC with an efficiency that can reach 60%.

5. Alkaline Fuel Cell (AFC): AFC uses alkaline potassium hydroxide as the electrolyte. These cells can achieve power generation efficiencies up to 70%.

PEMFC and SOFC are the most mature in Europe and also fit the best with the LEV requirement, so they have been selected. Fuel cells are not presently used for space applications in Europe. It is mandatory for the dedicated development and qualification phase to start soon for technology readiness in 2015. In Apollo, AFC type fuel cells were used. Using H2/O2 fuel cells has the extra benefit that water is the product of the reaction and can be used for life support. For computing the fuel cell’s architecture (oxygen and hydrogen storage, fuel cells mass), the model developed in RD[66] is used. For space applications, the volume of the fuel cell system is one of the most important design drivers. Therefore the storage of the fuels has to be optimised in term of volume. Both fuels (O2 and H2) can be stored in liquid form or gaseous form. Typically, to reduce the volume of the tanks, hydrogen is stored with a pressure range between 300 and 700 bars. However, this type of storage is not recommended for human missions. On the other hand, to store them in a liquid form without any loss by boiling off implies keeping the tanks at low temperatures.



page 175 of 317

s Storage Mass of storage for 1 kWh stored Hydrogen: metal 5.5Hydrogen: composite 0.79Hydrogen: carbon 0.22Hydrogen: carbon tube 1.1Hydrogen: liquid 0.28Oxygen: gaseous 1.76Oxygen: liquid 0.67

Table 6-48: Fuel storage options

All the possible storage options for the fuels are shown in Table 6-48. In fuel cell applications on Earth, the hydrogen is typically stored in gaseous form in composite cylinders. The other technologies (carbon nanofibres, carbon nanotubes and metal hybride) are not mature enough to be considered in this study (see RD[64] and RD[65] for more details). At ambient temperatures, the boil-off will be about:

• 0.3%/day for the hydrogen • 2.5%/day for the oxygen

6.13.2.2.1.2 Photovoltaic conversion

Different space-qualified photovoltaic technologies exist. GaAs triple junction cells have the best performances. Currently, in AM0 (28ºC) conditions, energy conversion of 27% can be reached. An increase in performance to 32% can be expected in 2015. The mass of a panel with these cells can be as low as 3.33 kg/m2, including the coverglass, the glue, the structure and the internal wiring. Several rigid panel architectures were considered in this study:

• Horizontal PVAs horizontal (either on the top of the vehicle or on the ground) • A plane PVA with a two-axis Sun-pointing device • Body-mounted PVA: some cells on the top of the vehicle and others on the vertical side

of the spacecraft (Figure 6-73.A) • Two opposite North-South tilted panels: (Figure 6-73.B) • Two plane PVAs mounted with a one-axis Sun pointing device: n-s or w-e directions

(Figure 6-73.C)



page 176 of 317

s

Solar Cells

2m

Ø=2.7m

North Direction

South Direction

Around 5m

1.5m

A) B)

C)

Figure 6-73: Body mounted and tilted panel options

Various developments are currently performed for optimising the solar panels’ weight by using flexible structures or thin films. In 2015, a mass of 0.53 kg/m2 can be expected with cells having an efficiency (AM0 28ºC) of about 15%. With that technology, the solar panels could consist of one or more blankets to unroll on the surface.

Triple junctions Thin Films

Table 6-49: Solar cells performances on the lunar surface



page 177 of 317

sDue to the uncertainty about the S/C attitude and the eclipses during descent and ascent, solar cells will only be considered for surface operations. To cover all landing dates and locations, a worst case of 90ºC is taken into account for the temperature of the photovoltaic cells. In addition to the losses shown in Table 6-49, 10% margin is also applied for taking into account the losses due to dust or from the dust removal mechanisms.

6.13.2.2.1.3 Solar dynamic

This option consists of a lens or a mirror that focuses the sunlight onto a receiver. In this receiver, the collected sunlight provides heat to a thermal-conversion unit. The NASA demonstrator has a low specific mass: 4.2 W/kg with 17% efficiency.

Figure 6-74: Solar dynamic demonstrators

This system is attractive as regards high power needs, but has the following disadvantages: • The need for a three-axis Sun-tracking system • The possible dust deposit of the mirror/lenses might considerably affect the performances

This concept has not been considered further.

6.13.2.2.1.4 Thermoelectric generators

The thermoelectric generator uses a thermal gradient to create electricity. In this case, the temperature difference between a point on the lunar surface exposed to the Sun and a point underground would be used. This source of energy is interesting when the solar illumination is too low to provide enough power. The efficiency of this system is low and the technology is still at conceptual level. This technology has not been considered further.

6.13.2.2.2 Secondary energy sources

A rechargeable storage system is required for the following reasons: • To supply the peak power on the bus (to complement the fuel cells or PVAs) • To provide additional power during the days with a low level of power production



page 178 of 317

s• Back-up source in case of failure of the power generation unit

Two types of technology are available: rechargeable batteries or regenerative fuel cells.

6.13.2.2.2.1 Rechargeable batteries

Compactness and low weight are the main drivers of the selection of a battery for this mission. Out of all the existing technologies (see Table 6-53), Li-Ion cells, in use onboard several European spacecraft, is selected. Performance is expected to reach in 2015 150 Wh/kg compared to the current 100 Wh/kg. The reduction of the electronic charge control, the optimisation of the packaging and modifications in the electrolytes are some of the identified means to reach this value.

Figure 6-75: Example of space-qualified Li-Ion battery module

6.13.2.2.2.2 Regenerative fuel cells

A description of this concept is in section 7.10.3.2.

6.13.3 Apollo design

The main difference with an Apollo mission is the duration of the stay on the lunar surface: 3 days versus 14 days. A pressurised and thermally controlled environment is mandatory involving a greater power requirement. Table 6-50 shows the main characteristics of the Apollo LEM. Due to the short descent, surface stay and ascent phases (in total less than four days), the LEM design could be based only on batteries.

Table 6-50: Apollo lunar module power subsystem budget

To optimise the mass of the Ascent Module, the batteries used during the descent and the surface operations are located in the part of the Descent Module that remains on the lunar surface.



page 179 of 317

sAssuming electrical performances for the AgZn batteries in line with the state of art at the time of the Apollo missions and the nominal capabilities of the LEM are:

• 44.8 kWh for the descent + surface stay • 8.3 kWh for the ascent phase

The mean total power consumption of the LEM on the surface was therefore less than 600W.

Table 6-51: Maximum power assessment of Apollo LEM based on the power design

Table 6-52: Power assessment of LEV based on the power budget

Table 6-51 and Table 6-52 compare the energy requirements of Apollo and the LEV mission of this study. At the time of the Apollo design, AgZn batteries were very attractive for short-duration missions, but presently the range of battery technologies available for space applications is enlarged by NiH2 and Li-Ion cells (see Table 6-53).

Battery Properties NiCd NiH2 Ag-Zn NiMH Pb - acid Li - Ion Lifetime (cycles at 80% DoD)

Moderate High Very Low Moderate Low Low

Watt - Hours/kg 35 65 100 55 20 125 Watt - Hours/Liter 85 80 185 180 200 318 Cell Voltage (V) 1.25 1.3 1.5 1.32 2.1 4.2 Discharge Rate High Moderate High High High High Memory Yes No No No No No Operating Temperature

0C – 50C 0C – 10C 0C – 45C 5C – 10C - 25C – 60C

Technology Maturity High Moderate Low Low High Moderate

Table 6-53: Current battery technology performances


Having identified all the different technologies, the corresponding power subsystem architectures have been computed to perform a trade-off. The power conversion and distribution units could



page 180 of 317

snot be properly traded in this feasibility study. In any case, the selection of the electronic topology is not a major parameter in this first study assessment. Standard parameters based on previous spacecraft are used and an average power conversion of 90% is also assumed.

Table 6-54: Options of the trade-off and key inputs for each phase

6.13.4.1 Descent and ascent phases

For both phases, the possible power sources are: 1. Fuel cells with an extra battery for power supply variations on the bus (able to

provide 500W during 6 hours) 2. Li-ion battery

Table 6-55: Mass budget in descent and ascent phases of the LEV power subsystem

Fuel cells are lighter and produces also around 17 litres of water that can be used by the life support. Therefore, they have been selected. Since the ascent phase might take place anytime during the surface stay, power source used for the descent phase cannot be recharged on the surface for use during the ascent. As a result, the vehicle will need to carry both systems. The total mass difference is then about 400 kg.

6.13.4.2 Lunar surface phase

The identified candidates are classified in two categories depending on their use of solar energy or not.

6.13.4.2.1 Architecture without solar cells

A system relying only on energy stored in a battery module would weigh more than 4 tonnes. A system relying on fuel cells (with a small battery for the peak power supply) would weigh 627 kg



page 181 of 317

s(752 kg with margin) and during a 14-day stay,would produce 193 litres of potable water. Since the life support has a requirement of 225 litres, the equivalent mass of the power subsystem for this trade-off is only 433 kg without margin.

Table 6-56: Power subsystem required for the surface stay when using fuel cells

6.13.4.2.2 Architecture with solar cells

Options with solar cells are more complex due to the constant variation of the Sun illumination and direction. Moreover, these parameters are directly linked with the latitude of the landing site. Each option includes photovoltaic cells for power generation and a power storage unit (either regenerative fuel cells or Li-Ion batteries).

6.13.4.2.2.1 Horizontally fixed solar arrays

For a horizontal PVA, not valid for polar locations, a 14-day surface stay can be divided into three consecutive periods (see Figure 6-76):

1. During the first days, the Sun is close to the horizon and the power generated is too low to provide the power required on the bus. The power storage unit is supplying the complementary power on the bus

2. Around the middle of the stay, the Sun’s direction is higher and enough power is generated for the bus users and for the recharge of the power storage unit

• During the last days, the Sun’s elevation is low again and the power is mainly provided by the power storage unit

•



page 182 of 317

s

Figure 6-76: Power balance during the surface operations

For high-latitude locations, horizontal solar arrays would have low performances due to the tilting of the sunrays on the panels. Consequently, only the locations in the latitude range [30ºS, 30ºN] are taken into account for this design. The selected method for sizing the solar arrays and the battery is based on an optimum use of these resources:

• The battery is fully discharged when power starts to be available on the bus (day 4 in Figure 6-76)

• The solar array is sized to provide the power generation required by this power balance method.

Three designs have been assessed:

Table 6-57: Options for the surface operations with horizontal solar panels

The computed mass budgets of these three designs are illustrated in Table 6-58. Shorter missions with a later arrival and an earlier departure would drastically decrease the requirement for the power storage. They have also been calculated. According to the results, a 14-day mission should be performed with regenerative fuel cells. For shorter missions (8-10 days), designs with battery modules become attractive compared to RFC module and could therefore be considered. These results also show that the reduced mass does not compensate the important increase of the area to be deployed in the thin-film option.



page 183 of 317

s

“Longest mission” best candidate+ Not dependant from the latitude!

PVA to accommodateBut not SADM-2 axis

“Short mission” best candidate

Depends of packaging

Huge area to deploy

Table 6-58: Assessment of the horizontal PVA options

6.13.4.2.2.2 Solar Array with a two-axis SADM

By mounting the PVAs on two-axis Sun-pointing mechanisms, the power generation is optimal and constant during the 14 days of surface operations (neglecting the albedo effect).

Light Solution+ Not dependant from the latitude!

PVA to accommodateBut not SADM-2 axis

Table 6-59: Assessment of the two-axis SADM PVA option

A battery module is also implemented for supplying the peaks on the bus. This module is recharged automatically during the low activity periods. The total mass of this design option is the lightest of all options considered and is applicable to the whole latitude range. The problem is the accommodation of the two-axis SADM that could not fit with the actual configuration of the LEV.

6.13.4.2.2.3 Body-mounted solar cells

A design with body-mounted solar cells was also been considered. It may lead to a low acceptable total mass and can cope with almost any landing site. However, because of the ascent



page 184 of 317

svehicle on top of the SHM, solar cells can be mounted only on the lateral surface. Consequently, this option has been rejected.

6.13.4.2.2.4 Two tilted solar arrays

Another concept, often found in the literature is based on two solar panels laying on the surface back to back with the same tilting angle but in opposite directions. Such a structure could be either mounted manually by the crew or being attached to the SHM and deployed automatically after the landing. Since the GNC system can perform an oriented landing, a defined orientation of this structure can be assumed. For any non-polar locations, the optimal attitude is to have the structure in the North-South direction. A first computation has shown that 61º tilted angle from the zenith direction is the optimal angle for the solar panels to have the highest local power generation during the 14 days on the lunar surface.

Efficiency compared to 2 axis SADM for a 30 latitude location

0

0,1

0,2

0,3

0,4

0,5

0,6

0 2 4 6 8 10 12 14

days

effic

ienc

y co

mpa

red

to o

ptio

n SA

DM

2-a

xis

Figure 6-77: Photovoltaic conversion efficiency of the tilted array option

The area required for the solar array is linked to the latitude of the module. For the latitude range [30N,30S], around 20 m2 of PVAs are required corresponding for example to two panels of five metres long and two metres high.

Table 6-60: Mass assessment of the 'tent' option in the [30ºN,30ºS] latitude range



page 185 of 317

s6.13.4.2.2.5 Solar array with a one-axis SADM

This option is a compromise between the solutions with fixed flat solar arrays with a limited latitude range and a PVA mounted to a complex two-axis Sun-pointing device. During the landing phase, the orientation of the SHM can be set with good accuracy. This initial positioning will be here used to compensate the loss of one degree of freedom. Rotation axis in the North-South (law 1) or in the East-West (law 2) direction has been considered. Both RFC and rechargeable battery modules are considered for the energy storage. Figure 6-78 and Figure 6-79 show the power balance evolution during the 14- day lunar surface stay with a battery storage.

Figure 6-78: One-axis SADM option: Lat 45 deg: Laws 1&2

Figure 6-79: One-axis SADM option: Lat 85 deg: Laws 1&2



page 186 of 317

sFor any latitude, with a solar panel correctly sized, the lunar stay can be divided into three successive phases for law 1:

1. Excess of power on the bus that is used for recharging the battery discharge during the descent

2. Lack of power on the bus that is compensated by the battery 3. Excess of power on the bus that can eventually be stored in the battery module

On the contrary, law 2 leads to a discharge-charge-discharge sequence. In law 1, by using the storage module of the DM, the additional storage requirement is small.

PVA Area requirement with a 200kg battery

0,0

5,0

10,0

15,0

20,0

25,0

30,0

-90 -60 -30 0 30 60 90

latitude

Are

a (m

2)

Law 1: Area RequiredLaw 2: Area Required

Law 2Law 1

Figure 6-80: Comparison area PVA required for both laws

Figure 6-80 shows (for a 200 kg battery module) the area of PVA required depending on the latitude for both pointing laws. By selecting a 15 m2 solar array, with the following laws, the power subsystem would be able to supply the bus requirements:

• Law 1 from latitude 60ºS to 75ºN • Law 2 for latitude lower than 60ºS

However, no landing higher than 75ºN can be achieved. Since the battery in the DM module is slightly bigger than 200 kg, the SHM power module does not need any additional power storage unit if law 1 is selected. As a result, for a landing between 75ºS and 75ºN, the mission can be fulfilled by having only 15 m2 solar array mounted to a 1-axis SADM in the Surface Habitation Module. The same computation was performed with RFC instead of battery for all LEV mission phases. In this case, the fuel required during descent is not sufficient for surface operations. As a consequence, the tank for the descent phase is increased. Table 6-61 shows the mass budget for the two storage options. The descent vehicle is strictly free of any power subsystem module. To minimise the mass of the ascent vehicle, only the required subsystems are implemented in it, the liquefaction units, the oxygen and hydrogen driers are located inside the SHM.



page 187 of 317

sThe advantage of using RFC instead of battery is large enough to justify development of the regenerative fuel cells technology for this application and also the development of a small module processing the liquefaction of the oxygen and hydrogen.

Table 6-61: Mass budget for one-axis SADM option with RFC and battery

6.13.4.2.3 Results

Table 6-62 shows a computation of the main benefits and disadvantages of each architecture for the power subsystem.



page 188 of 317

s

• Short mission scenario: ++• 14 days mission scenario: - • technology exists and qualified: (only improvements expected) • Not adapted for higher latitudes

• All mission scenarios: ++• Technology exists and qualified(only improvements expected)• more flexibility for landing latitudes•Accomodation of SADM: --

• Short mission scenario: ++• 14 days mission scenario: -- • blanket thin films accomodation: ?? (Not a mature european techno) • might become interesting depending of techno/accomodation impro vements• Not adapted for higher latitudes

• All mission scenarios: +• technology new in space• interesting in case of night survival mission requirements •Not adapted for higher latitudes

• Definitively too heavy

• All mission scenarios: ++ • Mass: - • Not affected by dust / landing attitude / local environment • Fuel Cells can be common with Descent/ Ascent Module • interesting techno also for HMM • Shareof Oxygen tank + provide water at end of mission

Solutions With Power Generation

Solutions Without Power Generation

• mass: ++ • deployment of the structure: -- • accomodation: --

• mass: +• All mission scenarios: ++• no mechanisms• dust deposit during landing: --•Accomodation: --•Distance from surface operations: -

•Latitude range: ++•Accomodation: ++•Qualification status: ++•Mass (SHM alone): +

• Latitude range: ++ • Accomodation: + • Qualification Status: --• Mass (SHM alone): ++ • Mass (LEV): ++

Table 6-62: Trade-off for power supply on the lunar surface

The final design selected for the LEV is regenerative fuel cells with a solar array mounted on a one-axis SADM.



page 189 of 317


Table 6-63: List of equipments for the LEV power subsystem

6.13.6 Option: dust removal mechanisms for solar arrays

To limit the degradation of the performances of the solar cells on the lunar surface, an option is to implement mechanisms on the PVAs to remove the dust. The advantages of such a system are:

• Limitation of the uncertainties of the dust deposit on the lunar surface • Decrease of the area of solar cells required • Less constraint for the accommodation of solar arrays (can be closer from the astronaut

activities location) The main inconvenience is that currently a space qualified mechanism does not exist. The following removal methods are considered:

• Mechanical excitation of the solar cells. Poor performances are expected in low-gravity atmosphere and for fine and electrostatic particles. Moreover, the accommodation is heavy and complex

• Wiping. the effectiveness and the damage of the coverglass are the critical issues of this concept

• Electrostatic repulsion method • Blowing with stored gas • Use of a transparent cover on top of the solar arrays. In this case, over-sizing of the solar

panels is anyway required due to the attenuation caused by the cover on sunlight



page 190 of 317

s In case the astronauts’ capabilities are used, the following should also be considered:

• Accessibility of the PVAs • Risk of damage to the photovoltaic panels, especially if contact is required

Figure 6-81: Possible dust removal mechanisms for solar arrays

No dust removal method has been selected in the baseline design.

6.14 LEV – thermal 6.14.1 Requirements and design drivers

The following set of main system requirements has been used in the thermal design of the LEV: • The thermal design shall cope with any lunar latitude. • The thermal design shall cope with a Moon-day permanence on the surface (up to 14

Earth days) including local noon. • The thermal design shall provide heat rejection from three crewmembers (on Ascent

Vehicle during descent and ascent, on SHM during surface stay). • The thermal design shall cope with an internal equipment max power dissipation of 1.5

kW. The main consequence of the first two requirements is that the thermal control subsystem of the LEV shall be sized taking into account a large range of sink temperatures. Figure 6-82 shows the evolution of temperature of the lunar surface as a function of the local time and of the latitude for the sub-solar longitude, as from a simplified thermal model. During one lunar day at equatorial sites the temperature may vary from about 120K (at dawn) to more than 350K (at noon).



page 191 of 317

s

Figure 6-82: Moon surface thermal model

A consequence of this is that significant IR heat fluxes shall be considered in the thermal balance of the vehicle and any radiator for heat rejection from the vehicle will have to be shielded from the surface.


Within this study, a trade-off has been performed between several thermal control concepts for the Surface Habitation Module:

1. Apollo LEM-type thermal design (active cooling + sublimator) 2. Active cooling + deployable radiator opportunely shielded from the surface to reduce the

sink temperature 3. Active cooling + the use of the subsurface as sink 4. Heat rejection based on heat pump principle

The latter technology has been discarded due to the present very low maturity for space applications. The lunar subsurface provides a natural sink, since below a few centimetres the temperature remains practically stable and low. However, due to the large distance from the surface of the heat sources, the conductive link to ground is little even when heat transfer through the legs is considered. This approach is therefore considered not practical.

Moon surface temperature in a day(t=0 at local dawn)

0.00

50.00

100.00

150.00

200.00

250.00

300.00

350.00

400.00

0 5 10 15 20 25

time (Earth days)

tem

pera

ture

(K)

Lat=0lat=40lat=80lat=88



page 192 of 317

sThe Apollo LEM design is mass efficient only for short surface durations and away from noon. The application of this design concept to the present mission case led to a sublimating mass required to cool the SHM in excess of 1 tonne, making it a poor competitor with the shielded radiator option. Option 2 has been retained as the only possible alternative to cope with the stated requirements. This design concept presents anyway several criticalities:

• The concept is unproven to date. • The radiator features anyway a large surface area and complexity due to the shielding and

the need of deployment mechanism. • Some amount of sublimating material is needed anyway for safety in case of non-prompt

deployment of the radiator at landing • Degradation of surface shield optical properties due to dust is critical

A significant development and testing effort is expected for qualification of this concept.


In order to achieve mass efficiency and taking into account the different durations in the different mission phases and the different vehicle functions, different thermal control designs for the LEV modules have been proposed. In the case of the DM only passive design with MLI covering tanks and structures and high temperature insulation for the main engine is required. Some heating power (100 W) is needed for thrusters and tanks to cope with a “cold” spacecraft attitude during cruise to the Moon. Eclipse when in Moon orbit has not been considered sizing because during this time the LEV receive power from the Hub. For the Ascent Vehicle an Apollo-like design with active cooling and sublimator has been selected. This is because the vehicle is sized for 1-day maximum operations and because any deployable radiator will not be consistent with the module mission (relatively high accelerations during flight) A schematic of the proposed system is shown in Figure 6-83:



page 193 of 317

s

Figure 6-83: Ascent vehicle thermal control system

The design features two active closed loops: primary coolant loop and secondary coolant backup loop with a water glycol solution. Each loop is split in two parts: a cabin loop and an equipment loop due to the different temperature requirements. The loops connected via heat exchangers. The batteries and electronic equipment are mounted on cold plates and rails through which coolant is continuously routed to remove waste heat. In-flight waste heat rejection from coolant loops is achieved by the water sublimators, which are sized for 1-day performance and are vented overboard. In case of temporary malfunction of the SHM system the Ascent Vehicle thermal control can be used as back-up for a few hours. Longer failures will require surface mission abort. High insulation is required to reduce heat from the environment in hot and cold cases. The SHM design is based on single phase, active cooling with a layout similar to the one of the Ascent Vehicle but with heat rejection provided by a deployable-shielded radiator (see Figure 6-84). In this configuration, the radiator is placed perpendicular to the surface and protected by the surface by means of a parabolic shade. The incident solar irradiation hitting the upper shade surface is reflected to a focal line above the radiator minimising heat leaks, while IR fluxes from the surface and possibly from vertical surfaces are deflected away by the underside of the shade. Specular solar reflectivity of the internal shade surface is fundamental (95% assumed in this study) while the external finishing shall minimise IR absorption. The design of the radiator is modular, meaning that the number of panels can be dimensioned according to the landing latitude with minor impact on the design.



page 194 of 317

sWithin this study a detailed analysis of the shielding has not been performed in all possible cases, a worst case “shielding efficiency” of 60% has been assumed in the final dimensioning.

Figure 6-84: Deployable radiator for the SHM

The worst-case analysis has shown that a surface area of 10 m2 is needed to cope with the equatorial landings. This implies the use of multiple (4) deployable panels (as shown in Figure 6-84), which can be folded together and later rolled down. The deployable radiator has been placed high up into the SHM to further reduce the view factor to the surface.


Table 6-64: DM thermal equipment list



page 195 of 317

s

Table 6-65: SHM thermal equipment list

Table 6-66: Ascent Vehicle thermal equipment list



page 196 of 317

s



page 197 of 317

s7 ORBITAL HABITATION VEHICLE (HUB)

7.1 Hub - systems The Hub is the orbital infrastructure that will be placed in an orbit around the Moon to fulfil the following functions:

• Provide a habitable volume for the astronauts, nominally six astronauts during six months at a time

• Provide safe haven for the surface crew in case of a solar storm or any other circumstances leading to surface mission abort

• Serve as docking station for the exploration vehicles As demonstrated in chapter 4 Mission architectures, Hub around the Moon increases the landing rate and provides a safer scenario for the astronauts.

7.1.1 Systems requirements

The system requirements for the Hub are shown in Table 7-1:

Table 7-1: Hub requirements



page 198 of 317

s7.1.2 System design drivers

An important design driver is the overall habitable volume required. Based on Figure 7-1, a crew of six for six months will require a total pressurised volume of 450 m3. Taking into account the constraint on the diameter of the launcher fairing, 4.5 m, this implies a cylinder of 28 metres height, which would have to be split into three modules to be launched. Therefore, an inflatable structure solution was adopted to allow a configuration of only two modules and reduce the number of launchers The launcher performance of 27 tonnes to LEO did impose the limit mass for each module. In case of an abort from the surface the orbital planes of the Hub and the ascent vehicle will not be the same, so a node change manoeuvre will have to be done. Instead of changing the node of the Hub (which will require an extra propulsion module) or the one of the ascent vehicle (which will severely oversize the LEV), it was found to be more practical to use a second LEV docked to the Hub as lifeboat. In case of abort from the surface, the second LEV would undock from the Hub, reach for the Ascent Vehicle, dock with it, allow for crew transfer and perform a second node change manoeuvre for redocking to the Hub. The total number of docking ports required will be four: two for the LEVs, one for the CTV or Cargo vehicle, and a spare. This is a major driver for the Hub configuration. Due to the relatively short stay duration of the crew on board (6 months), it was not required to include a short arm centrifuge in the Hub. For the orbit selected for the Hub (a 100 km altitude polar orbit), maintenance will be required, as it is unstable in terms of periselenium altitude. Therefore an AOCS system will have to be implemented. Although the mass of the Hub itself may not be large, several elements will be docked to the Hub during some of the time (LEVs, CTV, cargo vehicle), making the average mass of the complex over time about 110 tonnes. The required ∆V for AOCS was estimated to be 210 m/s per year, which over the mass of the complex leads to a propellant requirement of 9.3 tonnes per year. Therefore, a dedicated cargo vehicle per year will be required to replenish the propellant.



page 199 of 317


Figure 7-1: Hub configuration

The Hub is composed of two main modules, launched and sent to the Moon in separate missions: • Rigid module, similar to an ISS module. It contains most of the equipment required for

the Hub, plus three docking ports and an airlock for EVA excursions. • Inflatable module, providing most of the required habitable volume for crew quarters,

sleeping areas experimental racks toilets, etc, and an extra docking port. Both of these modules will have their own AOCS subsystem and will be autonomous for a TBD period of time in case of failure of the other module (redundancy for the astronauts). The system has been designed as a whole, meaning that no split between the two modules has been done at subsystem level. This will have to be further analysed in the next steps of the project as it might have a non-negligible impact on the design.



page 200 of 317

s7.1.4 Mass budgets

The constraint on the total wet mass of the Hub was 2x27 tonnes. To use launcher capabilities effectively, once a dry mass was achieved, the propulsion system for AOCS was sized to carry the maximum quantity of propellant from the start. This strategy should be further analysed in the future together with the refuelling strategy.

Table 7-2: Hub mass budget

7.2 Hub - configuration During the study, the time available to study the Moon-orbiting Hub was not sufficient to fully describe the internal parts of the modules. The required volume is overall sufficient to accommodate all relevant equipment, life support and other elements for the mission.


For the configuration of the Hub, the following set of requirements and design drivers were taken into account:

• The selected launcher, Ariane-5, volumetric capability, which mainly drives the diameter to 4.57 metres

• The total pressurised volume required for the crew in the Moon-orbiting Hub is 450 m3

• The Hub shall have four docking ports and one airlock • Good clearance shall be provided for rendezvous and docking/berthing



page 201 of 317

s7.2.2 Assumptions

Due to some similarity between the presently studied mission to the Moon and the Human Mission to Mars, some features of the design for the Hub have been taken from the latter and slightly altered to fulfill the requirements. The following additional assumptions were made for the Hub:

• Habitation Modules (HMs) are launched separately with Ariane-5: one rigid and one inflatable module

• Two solar arrays and radiators are attached later when modules are in Earth orbit


The design driver for the configuration of the HM is on one side the required pressurised volume (habitability requirement) and on the other the constraint of the available volume inside the launcher fairing. It was therefore decided to choose, apart from a rigid HM, an inflatable HM to increase the available volume in-orbit (see Figure 7-2).

Rigid Habitation Module

Inflatable Habitation Module

Figure 7-2: Hub configuration

7.2.3.1 Rigid Habitation Module

The rigid HM provides, in addition to the overall volume, a dedicated part for increased radiation environments, where the crew can shelter (see Figure 7-3). This module has the following characteristics:

• Length 12.305 m • Outer diameter 4.48 m • Gross volume 154 m3, volume 35 cm wall thickness 52 m3, net volume 102 m3 • Two docking ports on both ends and three docking ports on the cylindrical wall • Four propulsion tanks with 1.16 m diameter and two pressure tanks with 0.65 m diameter

and the thrusters



page 202 of 317

s

Rigid Habitation Module

Radiation Shelter

Figure 7-3: Rigid Habitation Module

Figure 7-4 shows the main measurements of the rigid HM:

Figure 7-4: Rigid HM measurements

Figure 7-5 shows the rigid HM inside the Ariane-5 long fairing, interfacing with the 3936 launcher interface:

Figure 7-5: Rigid HM in Ariane-5 long fairing

7.2.3.2 Inflatable Habitation Module

The module has the following characteristics: • Length 11.343 m • Diameter ‘backbone’ cylinder 2.624 m • Diameter inflated part 8.3 m, length 8.5 m • Gross volume 446 m3, volume 30 cm wall thickness 93 m3, net volume 352 m3 • One docking port at one end and an airlock with a port on the cylindrical wall • Four propulsion tanks with 1.16 m diameter and two pressure tanks with 0.65 m diameter

and the thrusters



page 203 of 317

s

Figure 7-6: Inflatable Habitation Module

Figure 7-7 shows the main dimensions of the inflatable HM:

Figure 7-7: Inflatable HM dimensions (when inflated)



page 204 of 317

sFigure 7-8 shows the inflatable HM inside the Ariane-5 long fairing, in deflated configuration. In the case of the inflatable HM, the launcher interface is with the 2624 adapter.

Figure 7-8: Inflatable HM in Ariane-5 long fairing

7.2.4 Hub configuration including two LEVs

The overall Hub configuration in lunar orbit is shown in Figure 7-9, providing a total of 454 m3 of net pressurised volume:

Figure 7-9: Hub configuration including two LEVs



page 205 of 317

sSome of the main dimensions of the Hub including LEVs are shown in Figure 7-10:

Figure 7-10: Hub main dimensions

7.3 Hub - life support 7.3.1 Requirements and design drivers

Table 7-3 shows the list of requirements for the life support system of the Hub design. Using a semi-closed life support system is envisaged, so recycling technologies will be necessary. The last parameter in the table represents the number of extra days of consumables that are carried. The extra consumables indicate the critical mass for recycling and will allow for an eventual emergency situation in case the equipment fails to function correctly.

Number of crew 6 #Mission duration 180 daysDrinking Water Requirement 3.91 kg / CM-dHygiene Water Requirement 10 kg / CM-dTotal number of EVA sorties 6 #Number of crew per EVA 2 #Duration of EVA sortie 6 hoursAirlock Volume 4.25 m3

Habitat Volume 450 m3

Pressure 101.3 kPaAtmosphere O2 content 21 %Atmosphere N2 content 79 %Number of Modules 2 #Days on stock (consumables) 15 days

Mission Design Inputs

Table 7-3: Mission design inputs



page 206 of 317

s7.3.2 Assumptions and trade-offs

7.3.2.1 Crew model

The crew model is similar to the one presented in section 6.6. The difference is the schedule for the astronauts. Only EVA and non-EVA days are taken into account. Due to the low number of EVA sorties (once per month), the average astronaut consumes about 2 500 kcal / CM-d. This leads to different rates in the global reaction that takes place in the human body (see Table 7-4)

Food Oxygen Faeces CO2 Urea H2OC 1 0 1 1 1 0H 1.9514257 0 1.8891842 0 4 2O 0.6230896 2 0.3943751 2 1 1N 0.0483719 0 0.1103540 0 2 0

0.8609712 0.9365373 0.0377393 0.8044909 0.0187411 0.7669304alpha beta gamma delta epsilon phi

PROPOSED STOECHIMOETRICS (in mols / h)

Stoechio

Table 7-4: Human metabolic activity

7.3.2.2 Storage

The storage capabilities are calculated as in section 6. A 10% margin is added on top of the nominal needs. The quantity of O2 and H2O to be carried is reduced from the nominal needs due to recycling. In this way, the quantity of wastewater to be stored is reduced. The oxygen to be carried is given by the following equation:

EVAleakageMissionOdailyneedstockdailyneedstored LossesLossesTROTOO ++⋅−⋅+⋅= )1( 2

The equation of water has to take into account three different recycling efficiencies (grey water to potable water, yellow water to grey water, black water to grey water). The water that is not recycled has to be stored as wastewater. The following equations show the potable and wastewater storage capabilities needed:

MissiongreycondensatedailyrategreydailyrateyellowOHyellowdailyrateblackOHblackdailyrate

EVALossesstockdailyneedstored

TROHOHROHROH

OHTOHOH

⋅−⋅++⋅+⋅+

++⋅=

)1()2222(

222

,,,2,,2,

,

MissiongreycondensatedailyrategreydailyrateyellowOHyellowdailyrateblackOHblackdailyrate

yellowOHyellowdailyrateblackOHblackdailyratewaste T

ROHOHROHROH

ROHROHOH ⋅

⎪⎭

⎪⎬⎫

⎪⎩

⎪⎨⎧

−⋅++⋅+⋅+

+−⋅+−⋅=

)1()2222(

)1(2)1(22

,,,2,,2,

,2,,2,

7.3.2.3 Atmosphere selection

For the Hub, the ISS atmosphere has been selected (101.3 kPa, 21% O2/79% N2). Mainly because EVA operations are not so frequent, a pre-breathing protocol such as that of the Shuttle or ISS can be afforded without any impact on the activities to be performed. This also avoids any increased fire risk and allows staying at sea-level equivalent. Using two different pressures for the Hub and LEV will incur additional operational constraints and procedures that must be followed for crew transfer.

7.3.2.4 Cabin / atmosphere losses model

The atmosphere losses due to leakage are calculated as described in section 6.6.



page 207 of 317

s7.3.2.5 EVA operations

The EVA operations have an impact on both the consumables and the leakage of water and atmosphere. The calculation is the same as it is for the LEV. The number of EVA operations is reduced to once per month (contingency EVA only).

7.3.2.6 Trade-off open loop / semi-closed loop

The type of life support system selected for the Hub is a semi-closed physico-chemical life support system. The system recycles O2 from CO2, and water according to certain intended recycling efficiencies (see Table 7-5):

Grey Water to Potable Water 95 %Yellow Water to Grey Water 95 %Black Water to Grey Water 20 %Solid Organic Waste to Food 0 %Oxygen 95 %

Recycling efficiencies

Table 7-5: Intended recycling efficiencies

The advantage of the semi-closed loop vs. the open loop is shown in Table 7-6, where the differences in mass, volume and power are shown.

Open Closed Relative Equipment Mass 1235 6150 -50% Consumables mass 1738 4040 -23% Volume 50.7 17.6 -35% Power 1250 3380 270%

Table 7-6: Trade-off open loop / semi-closed loop

The penalty in power can be accepted as the reductions in mass and volume are very important.


Figure 7-11 shows a simplified schematic of the life support system on the Hub.



page 208 of 317

s

Figure 7-11: Simplified schematic of the Hub life support system


Table 7-7 shows the equipment (items and quantities) that have been selected for the design:



page 209 of 317

sElement 1 Unit Name

1 Heat Exchanger 4 19.9 Fully developed2 Water Separator 4 9.8 Fully developed3 Fan and motor 4 18.6 Fully developed4 Pump and motor 6 24.0 Fully developed5 ISS HEPA Bacteria Filter Assembly 37 6.0 Fully developed6 MCA (Major Constituents Analyzer) 1 54.0 To be modified7 ANITA 1 53.0 To be developed8 ISS Water Processor 1 476.0 To be modified9 ISS Urine Processor 1 128.0 To be modified10 ARES 2 202.0 To be developed11 DISHWASHER 1 40.0 To be developed12 Fluid interface connection 450 1.0 To be modified13 IMV Valve 1 5.3 Fully developed14 IMV Fan 1 4.7 Fully developed15 Manual Pressure Equalization Valve 2 1.2 Fully developed16 Vent / Relief Valve 6 5.4 Fully developed17 Potable Water Gun 2 5.0 To be developed18 Potable Water Heater 1 5.0 To be developed19 Commode/Urinal 1 50.0 To be developed20 Personal Hygiene kit 6 1.5 To be developed21 Medical and Exercise Subsystem 1 592.3 Fully developed22 Smoke detectors 33 1.5 Fully developed23 Portable Fire Extinguishers 4 7.8 Fully developed24 APOLLO Suits 6 15.5 Fully developed25 Clothes 30 5.0 To be developed26 WASHER / DRYER 1 80.0 To be developed27 SHUTTLE PLSS 3 123.0 Fully developed28 Water Pressure Regulator 4 0.4 Fully developed29 Oxygen Pressure Regulator 8 2.0 Fully developed30 Nitrogen Pressure Regulator 8 2.0 Fully developed31 Filter 16 0.1 To be developed32 Transducer and switch 4 0.5 Fully developed33 Oxygen Valve 12 2.0 Fully developed34 Nitrogen Valve 12 2.0 Fully developed35 Water Valve 30 1.8 Fully developed36 Potable Water Tank 2 298.0 To be modified37 Waste Water Tank 2 174.6 To be modified38 Oxygen Tank 2 76.3 To be modified39 Nitrogen Tank 2 154.4 To be modified40 LEB Food Storage Box Descent 97 5.0 Fully developed41 LEB Food Storage Box Descent 59 5.0 Fully developed42 HEPA Filter 74 2.0 Fully developed43 MIDASS 1 28.0 To be developed

43 6149.9

Mass per Maturity LevelMASS [kg]Element 1: Habitation Module

ELEMENT 1 SUBSYSTEM TOTAL

Unit Quantity

Table 7-7: Hub life support system equipment summary

The equipment used has been selected by choosing the optimum components from a database including ISS, Shuttle, Spacelab, CRV, Apollo and Soyuz data, as well as some advanced foreseen technologies. The mass and power data of the different units are described in RD[25], RD[26] and RD[27]. The overall Life Support System mass and power budget is shown in Table 7-8.



page 210 of 317

sOutput Hub Vehicle

Total Life Support Mass [kg] 10193.93 - Total Equipment Mass [kg] 6149.85 Margin [%] 0.09 - Total Consumables Mass [kg] 4043.99 Oxygen [kg] 260.06 Potable Water [kg] 2257.92 Hygiene Water [kg] 0 Food (dry) [kg] 594.68 Food Packaging [kg] 297.34 Waste solid allowance [kg] 108 Nitrogen [kg] 525.99 Total Power [W] 3384.44

Table 7-8: Mass and power budgets of the Hub life support system

7.4 Hub - propulsion This section describes the Hub’s integrated storable propulsion system that is required to perform attitude and orbit control for the lifetime of the Hub in LLO and the propulsion modules required for the transfer of the Hub, Crew Transfer Vehicle (CTV) and Lunar Excursion Vehicles (LEV) to the Moon and back (only for the CTV). To deliver each payload module to the Moon, the payload limitations of the Ariane-5 determine the use of multiple stacked propulsion systems. Table 7-9 shows the propulsion modules to be used for each single mission, as well as the required number and the total mass and type of propulsion stages required to perform a given manoeuvre for that specific payload. For example, the final propulsion system configuration for a CTV mission will consist of two cryogenic propulsion systems each of 27 tonnes, fired separately, to achieve Trans Lunar Injection (TLI); a single cryogenic propulsion module to inject the payload into lunar orbit, Lunar Orbit Insertion (LOI); and an additional storable propulsion module to inject the CTV into Trans Earth Injection (TEI).

To be Transferred Manoeuvre Stage Req. Propellant No. of Stages

TLI 27 tonne Cryogenic 2 27 tonne Hub (1 & 2) LOI 8.6 tonne Cryogenic 1 TLI 27 tonne Cryogenic 2 LOI 10 tonne Cryogenic 1 CTV TEI 10.6 tonne Storable 1 TLI 27 tonne Cryogenic 2 LEV, DM LOI 10 tonne Cryogenic 1 TLI 16 tonne Cryogenic 2 LEV, SHM & AV LOI 7 tonne Cryogenic 1

Table 7-9: Earth-Moon transfer propulsion systems



page 211 of 317

sFor manufacturing purposes and to minimise inherent costs, a common propulsion system design is taken for each of the cryogenic propulsion system stages. As shown in Table 7-9, different sizes of both the cryogenic and the storable stages will be required. For the purposes of the study, three design points have been defined: 27 tonnes and 10 tonnes for the cryogenic stage and 10.6 tonnes for the storable one. A model was developed for their design at system level. The model takes into account the tanks’ design, structures, engines and pipework and preliminary thermal insulation and debris shielding. Other subsystems were not taken into account, but their impact was assumed to be low.

7.4.1 Habitation module integrated propulsion system

7.4.1.1 Requirements and design drivers

The Hub’s propulsion system is required to perform a velocity increment of ∆V = 210 m/s per year for AOCS, using storable propellant. Storable propellant is selected because the lifetime in LLO is long, and the thrust required is low enough. Throughout the mission lifetime there will be different numbers of elements attached to the Hub, so, the overall system mass in LLO is a function of time. An average total wet mass of 110.5 tonnes has been taken to size the propulsion system. As the requirement is only to provide attitude control capability, so the thrust required is low and therefore a pressure-fed system can be used. The pressure-fed system enables greater system reliability over a pump-fed system, and can be operated at higher pulse mode frequencies for attitude-thrusting manoeuvres.


Two identical propulsion systems are designed for each Hub module. Each of them is based on a MON/MMH bipropellant propulsion system with 16 x S400-2 EADS thrusters. These are 400 N thrusters and fit on the exterior of each Hub module for launch. Each Hub propulsion system is designed to support a dry ‘payload’ mass of 50950 kg (as computed over the mission lifetime), and capable of producing a total velocity increment of ∆V = 210m/s. Four Titanium-Alloy propellant tanks contain the required fuel and oxidiser and two pressurant tanks contain helium gas to maintain propellant tank pressure at a nominal value of 22 bar. The design of the propulsion system is shown in Figure 7-12. The separate integrated propulsion systems for each Hub module are shown. The propellant and pressurant tanks for both modules are situated on the exterior of both Hub modules surrounding the end docking ports:



page 212 of 317

s

Figure 7-12: AOCS propulsion modules configuration

Figure 7-13 shows the 400 N engine and Table 7-10 shows the main S400-2 engine characteristics.

Figure 7-13: S400-2 EADS thruster

A nominal value of 318s is taken as baseline specific impulse in the preliminary design of the propulsion system.

Propellant and

Pressurant Tanks on

Propellant and

Pressurant Tanks



page 213 of 317

sValue Unit

Manufacturer EADSCycle Pressure-fedVacuum Thrust 400 NSpecific Impulse 318 sChamber Pressure 10 barPropellants MMH/NTOOverall Length 0.531 mNozzle exit diameter 0.248 mTotal Mass 3.6 kg Table 7-10: S400-2 EADS engine characteristics

Four propellant tanks are required. The tanks are bolted to shell structure around an equatorial support ring. Titanium is used for propellant compatibility and high strength-to-density ratio. The tanks are stabilised against buckling by maintaining an internal pressure internal nominal pressure of 22 bars. This pressure is the nominal pressure level required for this system design to ensure a 10 bar chamber pressure in the thrusters combustion chambers. The numerically estimated tank mass includes a Propellant Management Device (PMD) in the design. Table 7-11 shows the tank characteristics.

Value UnitPropellant MMH/NTOMEOP 2.5 MPaTank Material Ti-6Al-4VVolume 1.19 m3

Diameter 1.32 mMinimum Wall Thickness 2.13 mmTank Mass 88 kgQuantity 4

Table 7-11: Hub module propellant tank characteristics

As regards pressurant tanks, the 22 bar nominal storage pressure is maintained through a helium pressurant system. The helium pressurant is stored in two Composite Over-wrapped Pressure Vessels (COPVs) for each Hub module propulsion system. The COPV consists of an inner titanium liner with graphite epoxy over-wrap, enabling considerable mass reduction. Table 7-12 shows the baseline pressurant tank characteristics.

Value UnitPropellant HeliumMEOP 31 MPaTank Material COPVVolume 0.211 m3

Diameter 0.74 mMinimum Wall Thickness - mmTank Mass 31 kgQuantity 2 Table 7-12: Hub module pressurant tank characteristics



page 214 of 317

sAs regards the Hub’s module propulsion system mass budget, Table 7-13 shows the ratio of dry mass to wet mass for each Hub storable propulsion module. Here the dry mass includes the pressurant mass, as well as a 10% subsystem margin.

Value UnitDry Mass 567 kgPropellant Mass 5552 kgTotal Wet Mass 6138 kg

Table 7-13: Hub module propulsion system mass budget

7.4.1.3 List of equipment

Table 7-14 shows the equipment summary and mass budget of the Hub’s storable bipropellant propulsion system. The mass budget includes a 10% subsystem dry mass margin and provides the pressurant mass separately.

Component Mass [kg]/unit Qty/Hub Module Engine 3.6 16 Line Filter 0.23 8 Latching Valves: Monopropellant 0.6 4 Bi-propellant 0.5 2 Pipe work 25 1 Pressure Transducer 0.4 7 Pressure Regulator 1.5 2 (Pyro) Normally Closed Valves 0.35 6 (Pyro) Normally Open Valves 0.35 4 Isolation Valves 0.5 4 Fill/Drain Valve/TP 0.1 14 Non-return Valve 0.2 4 Propellant tank (ox shell) 60 2 Propellant tank (fu shell) 61 2 Pressurant tank 21 2 Flight Control Orifice (fuel and oxidiser) 0.1 4 Dry Mass Margin 10 %

Propulsion System Dry Mass 567 Propellant Mass 5552

Pressurant Mass 18 Propulsion System Wet Mass 6138

Table 7-14: Hub module’s bipropellant propulsion system equipment and mass budget

7.4.2 Translunar injection propulsion system

This propulsion system is required to place a vehicle from LEO into TLI.



page 215 of 317

s7.4.2.1 Requirements and design drivers

The 27 tonne total wet mass launch limitation is the baseline design driver. The final configuration must fit within the given envelope of the Ariane-5 fairing. With short mission duration, and a desirable high specific impulse, a cryogenic propulsion system is chosen. To minimise gravity losses during thrusting manoeuvres, a thrust in the range of 65 kN is required. To achieve the correct trajectory during engine firing, a gimbal mount is required for thrust vector control. A Reaction Control System (RCS) is also required for attitude control. The RCS is required to perform a velocity increment of ∆V = 100 m/s. Insulation is required around the cryogenic propellants tanks to minimise boil-off of the cryogenic propellants. The propulsion system will also require shielding from potential meteorite impact, as it will spend a few months in LEO during the assembly.

7.4.2.2 Assumptions and trade-offs

For this 65 kN thrust application, a turbopump-fed system is used to achieve the necessary engine mass flow rate and chamber pressure while keeping a relatively low storage tank pressure. A reaction control system is required for attitude control, and a separate storable pressure-fed system, in a suitable thrust range for reaction control, is chosen.


From the 27-tonne total wet mass baseline design requirement, an iterative design process is taken to calculate the final ratio of propulsion dry mass to total system mass. The propulsion system consists of a cryogenic main engine fuelled with hydrogen and oxygen, and has a separate RCS, which uses storable propellants. The baseline design configuration and the propulsion stage main characteristics are outlined below. As regards configuration, Figure 7-14 shows the general TLI stage propulsion system. The liquid oxygen tank is situated at the base of the common bulkhead cryogenic propellant tank for structural and dynamic stability during launch. Four helium tanks pressurise the main cryogenic tank and are situated beneath the cryogenic tanks thrust frame assembly. The reaction control system is situated above the main cryogenic stage, with thrusters on the exterior for attitude control manoeuvres. A ringed adapter extends sufficiently from each end of the cryogenic tank such that the propulsion systems can be stacked in series.



page 216 of 317

s

Figure 7-14: TLI propulsion system configuration

As regards the cryogenic stage main engine, the HM7-B is chosen as baseline. The HM7-B has much flight heritage, having been used on Ariane 1-4 launch vehicles, and now in Ariane-5 in the cryogenic upper stage. The engine is only required for a single firing. The HM7-B is in Figure 7-15.

Figure 7-15: HM7-B cryogenic engine



page 217 of 317

sTable 7-15 shows the HM7-B’s main engine characteristics. The engine will have gimbal capability for thrust vector control and will be attached to the cryogenic tank through a thrust frame adapter; whose respective estimated masses are given in the equipment summary.

Value UnitManufacturer Snecma MoteursCycle Gas GeneneratorVacuum Thrust 64.8 kN Specific Impulse 446 sCombustion Pressure 37 barArea ratio 83.1Propellants LOX-LHPropellant flow rate 14.8 kg/sMixture ratio 5Turbine speed 60,800 rpmTurbine power 400 kW Height 2.01 mNozzle exit diameter 0.99 mTotal weight 165 kg

Table 7-15: HM7-B engine characteristics

As regards propellant tanks, to maximise the available envelope volume of the Ariane-5 launch vehicle fairing, a single cryogenic tank was chosen. Thus the cryogenic propulsion system configuration is effectively a booster stage system. The tanks are arranged in tandem for minimal overall weight and maximum storage volume in the given envelope. Table 7-16 shows the hydrogen and oxygen liquid properties and storage conditions used in this preliminary design:

Value UnitLOX storage temperature 90 KLOX vapour pressure 1.00E+05 PaLOX density 1142 kg/m3

LH storage temperature 20 KLH vapour pressure 1.00E+05 PaLH density 71 kg/m3

Table 7-16: Cryogenic propellant properties and storage conditions

Table 7-17 shows the properties of the aluminium-lithium alloy material used as the cryogenic tank shell material:

Value UnitMaterial Aluminium-LithiumDensity = 2540 kg/m3Ultimate Stress= 470 MPaYield Strength= 330 MPaYoungs Modulus= 79 GPaTemperature = 298.16 KUltimate strength/safety factor 235 MPa

Table 7-17: Cryogenic propellant tank materials



page 218 of 317

sA safety factor of two is assumed which results in a maximum allowable operational stress of 235 MPa. Table 7-18 shows the liquid oxygen cryogenic tank dimensions:

Value UnitTotal Volume 16.8 m3Maximum Diameter 3.52 m

a 1.76 mb 0.98 mk 1.8Volume Ellipsoidal End 6.3 m3Volume Cylinder 4.1 m3Height Cylinder 0.43 mTotal Tank Height 2.38 mThickness Ellipsoid End 2.9 mmThickness Cylinder 3.4 mmMass Ellipsoidal End 16.0 kgMass Cylinder 40.9 kgTank Mass 72.8 kgTank Mass + 30% tank factor= 94.7 kg

Table 7-18: TLI liquid oxygen tank characteristics and dimensions

In this table, a = elliptical tank-end major half diameter and also the radius of the cylindrical tank section; b = elliptical tank-end minor half-diameter; k is the tank-end ellipse ratio=a/b, where k = 1 for a spherical end. The tank shell structure must be able to support the maximum expected internal tank pressure, which is a function of the necessary storage liquid vapour pressure, the expected acceleration in flight, and the corresponding additional hydrostatic pressure. Table 7-19 shows these factors and provides the necessary design tank storage pressure.

Value UnitMax. acceleration 49 m/s2Estimated Tank Height 2.38 mLOX vapour pressure 100000 PaLOX hydrostatic pressure 133197 PaPressure Loss in system 60000 PaOxidiser tank pressure 293197 Pa Table 7-19: TLI liquid oxygen tank pressure properties

The LOX tank thus has an internal design tank pressure of approximately 3 bars. A 3-4 mm aluminium-lithium alloy wall thickness supports this pressure. Table 7-20 shows the dimensions for the liquid hydrogen cryogenic tank. The thickness does not include the additional thickness resulting from thermal insulation and shielding from meteorite impact.



page 219 of 317

sValue Unit

Total Volume 70.7 m3Maximum Diameter 3.5 m

a 1.76 mb 0.98 mk 1.8Volume Ellipsoidal End 6.3 m3Volume Cylinder 58.1 m3Height Cylinder 6.0 mTotal Tank Height 7.9 mThickness Ellipsoid End 1.85 mmThickness Cylinder 1.76 mmMass Ellipsoidal End 10.2 kgMass Cylinder 294.4 kgTank Mass 314.8 kgTank Mass + 30% tank factor= 393.6 kg

Table 7-20: TLI liquid hydrogen tank characteristics and dimensions

The LH internal tank pressure has a design pressure of approximately 2 bar (Table 7-21). A 1-2 mm aluminium-lithium wall thickness supports this pressure. Thus, the bulkhead common to both tanks experiences a small differential pressure.

Value UnitMax. acceleration 49 m/s2Estimated Tank Height 7.94 mLH vapour pressure 100000 PaLH hydrostatic pressure 27640 PaPressure Loss in system 60000 PaFuel tank pressure 187640 Pa

Table 7-21: TLI liquid hydrogen tank pressure properties

The internal pressure is maintained through a pressurant system, as shown in Table 7-22, to help stabilise the cryogenic tank structure. The pressurant system consists of four externally mounted composite over wrapped pressure vessels.

Value UnitPropellant HeliumMEOP 31 MPaTank Material COPVVolume 0.26 m3Diameter 0.79 m Minimum Wall Thickness - mmTank Mass 41.0 kgQuantity 4

Table 7-22: TLI cryogenic propulsion system pressurant tank characteristics



page 220 of 317

sThe tank structure is reinforced with additional self-supporting structures (stringers). Thermal insulation is required to minimise boil-off of propellant from the cryogenic tank and consists of multiple laminate layers of Kapton, separated by Dacron Net (see Table 7-23):

Value UnitThickness 3.81E-06 mFuel Tank surface area 80.85 m2No. Layers 320Density Kapton 0.011 kg/m2Density Dacron net 0.00541 kg/m2Density Betacloth 0.26 kg/m2Mass of MLI 445.1 kg

Table 7-23: TLI cryogenic propellant tank multi layer thermal insulation

Note that the thermal insulation is simply a preliminary estimation to obtain a more accurate mass budget of the propulsion system. The amount of propellant lost through boil-off from the cryogenic tanks during storage in LEO and during the LTO has been taken into account at system level. As regards the cryogenic propulsion system mass budget, the cryogenic stage characteristics are shown in Table 7-24. The RCS is excluded from this budget, and the dry mass includes a 10% subsystem margin as well as the helium pressurant required for the cryogenic stage.

Value UnitDry Mass (incl.pressurant) 2816 kgLiquid Oxygen Mass 19146 kgLiquid Hydrogen Mass 3829 kg

Table 7-24: TLI cryogenic stage mass budget

As regards reaction control system, an auxiliary RCS, using storable propellants, is used and required for fine attitude control. The RCS uses 16 European Apogee Motors (EAM), with MMH and NTO storable propellants, which are pressure-fed into each thruster’s combustion chamber. The status of the EAM is still in development and so the engine characteristics, as shown in Table 7-25, are used as preliminary design values:

Value UnitManufacturer EADSVacuum Thrust 500 NSpecific Impulse 325 sArea ratio 83.1Propellants MMH/NTOMixture ratio 1.63Height 0.7 mNozzle exit diameter 0.3 mTotal weight 5 kg

Table 7-25: EAM engine characteristics



page 221 of 317

sThe RCS is a pressure-fed system, consisting of two fuel and two oxidiser propellant tanks, and two helium pressurant tanks. The propellant masses and RCS dry mass are shown in Table 7-26, where the dry mass includes the mass of the helium pressurant and a 10% subsystem margin.

Value UnitDry Mass 225 kgMMH Mass 332 kgNTO Mass 542 kg

Table 7-26: TLI RCS mass budget

The preliminary design for the propellant tanks assumes Ti-6AL-4V as baseline material. The tanks are attached to the vehicle structure by equatorial rings. The numerically estimated tank mass includes a Propellant Management Device (PMD) in the design. Table 7-27 shows the tank characteristics.

Value UnitPropellant MMH/NTOMEOP 2.5 MPaTank Material Ti-6Al-4VVolume 0.19 m3

Diameter 0.71 mMinimum Wall Thickness 1.3 mmTank Mass 16.5 kgQuantity 4 Table 7-27: TLI RCS propellant tank characteristics

The helium pressurant is stored in two Composite Over-wrapped Pressure Vessels (COPVs). The pressure tanks have a nominal storage pressure of 250 bars. Table 7-28 shows the baseline pressurant tank characteristics for the TLI RCS propulsion system.

Value UnitPressurant HeliumMEOP 31 MPaTank Material COPVVolume 0.033 m3

Diameter 0.4 mMinimum Wall Thickness - mmTank Mass 6.1 kgQuantity 2

Table 7-28: TLI RCS pressurant tank characteristics

As regards the overall mass budget, Table 7-29 shows the main TLI propulsion system characteristics, where the dry mass includes a 10% subsystem margin as well as the pressurant required for both the cryogenic and storable systems, and where the propellant mass includes the RCS propellant.



page 222 of 317

sValue Unit

Dry Mass 3063 kgPropellant Mass 23937 kgTotal Wet Mass 27000 kg

Table 7-29: TLI propulsion system characteristics


As regards reaction control system, Table 7-30 shows the equipment summary and mass budget of the reaction control system for the TLI propulsion system. The mass budget includes a 10% subsystem dry mass margin and provides the pressurant mass separately.

Component Mass [kg]/unit Qty Main Engine 5 16 Line Filter 0.23 8 Latching Valves: Monopropellant 0.6 4 Bi-propellant 0.5 2 Pipe work 25 1 Pressure Transducer 0.4 7 Pressure Regulator 1.5 2 (Pyro) Normally Closed Valves 0.35 6 (Pyro) Normally Open Valves 0.35 4 Isolation Valves 0.5 4 Fill/Drain Valve/TP 0.1 14 Non-return Valve 0.2 4 Propellant tank (ox shell) 16.3 2 Propellant tank (fu shell) 16.5 2 Pressurant tank 6.1 2 Flight Control Orifice (fuel and oxidiser) 0.1 4 Dry Mass Margin= 10 %

Propulsion System Structural Mass 222 Propellant Mass 874


Table 7-30 TLI RCS equipment summary and mass budget

As regards translunar injection propulsion system, Table 7-31 shows the equipment summary and mass budget of the entire Translunar Injection Propulsion System. The mass budget includes a 10% subsystem dry mass margin and provides the pressurant mass separately.



page 223 of 317

sComponent Mass [kg]/unit Qty Main Engine (Translunar Injection Engine) 165 1 Line Filter 0.23 6 Latching Valves: Monopropellant 0.6 4 Thruster (bi-prop) 30 2 Pipe work 20 1 Pressure Transducer 0.4 8 Pressure Regulator 1.5 2 Fill/Drain Valve/TP_Pressurant 0.05 4 Fill/Drain Valve/TP_Propellant 6 4 Common Bulkhead Tank 933 1 Pressurant tank 41 4 Thrust Frame Assembly 373 1 Adapter 200 1 Meteorite Shielding 243 1 Gimbal system 327 1 RCS 222 1 Dry Mass Margin= 10 %



Table 7-31: TLI propulsion system equipment summary and mass budget

7.4.3 Lunar orbit insertion propulsion system

This propulsion system is required to place a given payload module from LTO into LLO by performing a LOI single thrust manoeuvre. The system is a downscale of the previously described 27-tonne cryogenic module. The changes carried out are related to the main tank length (lower mass of propellant required) and the resizing of the RCS system. Therefore only the summary tables are presented hereafter.



page 224 of 317

s Value Unit

Total Volume 5.89 m3 Maximum Diameter 3.52 m a 1.7575 m b 0.87875 m k 2 Volume Ellipsoidal End 5.68 m3 Volume Cylinder 5.89 m3 Height Cylinder 0.61 m Total Tank Height 1.49 m Thickness Ellipsoid End 2.50 mm Thickness Cylinder 2.84 mm Mass Ellipsoidal End 12.48 kg Mass Cylinder 48.39 kg Tank Mass 73.4 kg Tank Mass + 30% tank factor= 95.4 kg

Table 7-32: LOI liquid oxygen tank characteristics and dimensions

Value Unit Max. acceleration 49 m/s2 Estimated Tank Height 1.49 m LOX vapour pressure 100000 Pa LOX hydrostatic pressure 83199 Pa Pressure Loss in system 60000 Pa Oxidiser tank pressure 243199 Pa

Table 7-33: LOI liquid oxygen tank pressure properties



page 225 of 317

s Value Unit

Total Volume 24.84 m3 Maximum Diameter 3.52 m a 1.76 m b 0.88 m k 2.00 Volume Ellipsoidal End 5.68 m3 Volume Cylinder 13.47 m3 Height Cylinder 1.39 m Total Tank Height 3.15 m Thickness Ellipsoid End 1.76 mm Thickness Cylinder 1.60 mm Mass Ellipsoidal End 8.77 kg Mass Cylinder 62.23 kg Tank Mass 79.8 kg Tank Mass + 30% tank factor= 99.7 kg

Table 7-34: LOI liquid hydrogen tank characteristics and dimensions

Value Unit Max. acceleration 49 m/s2 Estimated Tank Height 3.15 m LH vapour pressure 100000 Pa LH hydrostatic pressure 10955 Pa Pressure Loss in system 60000 Pa Fuel tank pressure 170955 Pa

Table 7-35: LOI liquid hydrogen tank pressure properties

Value Unit

Propellant Helium MEOP 31 MPa Tank Material COPV Volume 0.075 m3 Diameter 0.52 m Minimum Wall Thickness - mm Tank Mass 13.1 kg Quantity 4

Table 7-36: LOI cryogenic propulsion system pressurant tank characteristics



page 226 of 317

s Value Unit Thickness 3.81E-06 m Fuel Tank surface area 28.89 m2 No. Layers 320 Density Kapton 0.011 kg/m2 Density Dacron Net 0.00541 kg/m2 Density Betacloth 0.26 kg/m2 Mass of MLI 159 kg

Table 7-37: LOI cryogenic propellant tank multi layer thermal insulation

Value Unit Dry Mass (incl.pressurant) 1375 kg Liquid Oxygen Mass 6727 kg Liquid Hydrogen Mass 1345 kg

Table 7-38: LOI cryogenic stage mass budget

Value Unit

Dry Mass (incl.pressurant) 180 kg MMH Mass 124 kg NTO Mass 200 kg

Table 7-39: LOI RCS mass budget

Value Unit Propellant MMH/NTO MEOP 2.5 MPa Tank Material Ti-6Al-4V Volume 0.07 m3 Diameter 0.51 m Minimum Wall Thickness 0.94 mm Tank Mass 7.99 kg Quantity 4

Table 7-40: LOI RCS propellant tank characteristics

Value Unit Propellant Helium MEOP 31 MPa Tank Material COPV Volume 0.012 m3 Diameter 0.29 m Minimum Wall Thickness - mm Tank Mass 3.20 kg Quantity 2

Table 7-41: LOI RCS pressurant tank characteristics



page 227 of 317

s Value Unit

Dry Mass (incl.pressurant) 1569 kg Propellant Mass 8431 kg Total Wet Mass 10000 kg

Table 7-42: LOI propulsion system characteristics

Component Mass [kg]/unit Qty Main Engine 5 16 Line Filter 0.23 8 Latching Valves: Monopropellant 0.6 4 Bi-propellant 0.5 2 Pipe work 25 1 Pressure Transducer 0.4 7 Pressure Regulator 1.5 2 (Pyro) Normally Closed Valves 0.35 6 (Pyro) Normally Open Valves 0.35 4 Isolation Valves 0.5 4 Fill/Drain Valve/TP 0.1 14 Non-return Valve 0.2 4 Propellant tank (ox shell) 7.9 2 Propellant tank (fu shell) 8 2 Pressurant tank 3.2 2 Flight Control Orifice (fuel and oxidiser) 0.1 4 Dry Mass Margin 10 %



Table 7-43: LOI RCS equipment summary and mass budget



page 228 of 317

sComponent Mass/unit (kg) Unit Main Engine (Lunar Orbit Insertion Engine) 165 1 Line Filter 0.23 6 Latching Valves: Monopropellant 0.6 4 Thruster (bi-propellant) 30 2 Pipe work 20 1 Pressure Transducer 0.4 8 Pressure Regulator 1.5 2 Fill/Drain Valve/TP_Pressurant 0.05 4 Fill/Drain Valve/TP_Propellant 6 4 Common Bulkhead Tank 354 1 Pressurant tank 13 4 Thrust Frame Assembly 142 1 Adapter 200 1 Meteorite Shielding 87 1 Gimbal system 124 1 RCS 179 1

Dry Mass Margin 10 %

Propulsion System Structural Mass= 1558

Propellant Mass = 8429 Pressurant Mass = 13

Propulsion System Wet Mass = 10000

Table 7-44: LOI propulsion system equipment summary and mass budget

7.4.4 Trans-Earth injection propulsion module

A propulsion module is also required to place the CTV into Trans-Earth Injection (TEI) from LLO.

7.4.4.1 Requirements and design drivers

The baseline requirement for this propulsion system is to achieve a total wet mass of 10.6 tonnes. The system requires that the propellant be storable to avoid boil-off. A thrust of 28 kN is required to minimise gravity losses. The final configuration must fit within the given envelope of the Ariane-5 fairing. To achieve the correct trajectory during engine firing, a gimbal mount is required for thrust vector control.

7.4.4.2 Assumptions and trade-offs

The Aestus engine is chosen to perform the Trans-Earth Injection manoeuvre. It uses storable bipropellant, is pressure-fed and meets the 28 kN thrust requirement. A reaction control system is required for attitude control, and as the main engine is pressure fed, a Unified Propulsion System (UPS) is chosen as the baseline design for the propulsion system.



page 229 of 317

s7.4.4.3 Baseline design

From the 10.6 tonne total wet mass baseline design requirement, an iterative design process is taken to calculate the final ratio of propulsion dry mass to total system mass.

Figure 7-16: TEI propulsion system configuration

As shown in Figure 7-16, the propellant is stored in four tanks and two helium pressurant tanks for the pressurant. The RCTs are mounted externally and provide all the required degrees of freedom for AOCS. As regards main engine, the Aestus (shown in Figure 7-16) is chosen as the baseline design engine because it meets the 28-kN thrust requirement. The Aestus has much flight heritage, powering the Ariane-5 standard version upper stage for the insertion of payloads into LEO, SSO and GTO.

Figure 7-17: Aestus engine

Table 7-45 shows the Aestus main engine’s characteristics. The engine will have gimbal capability for thrust vector control and will be attached to the propellant tank assembly through a thrust frame adapter, whose respective estimated masses are given in the equipment summary.



page 230 of 317

s Value Unit Manufacturer EADS Cycle Pressure-fed Vacuum Thrust 28 kN Specific Impulse 324 s Chamber Pressure 11 bar Area ratio 84 Propellants MMH/NTO Propellant flow rate 8.8 kg/s Overall Length 2.2 m Nozzle exit diameter 1.32 m Total Mass 111 kg

Table 7-45: Aestus engine’s characteristics

As regards propellant tanks, four are required. Each tank has an individually welded, cylindrical mid-section with cassini domes, each made from titanium alloy (see Figure 7-18).

Figure 7-18: Titanium propellant tank

The tanks are bolted to shell structure around an externally mounted support ring. The tanks are stabilised against buckling by maintaining an internal nominal pressure of 22 bars. This pressure is required to ensure an 11-bar chamber pressure in the main engine and the RCT. The numerically estimated tank mass includes a Propellant Management Device (PMD) in the design. Table 7-46 shows the propellant tank characteristics. The tank diameter is 1.56 m, which is the maximum allowable diameter for four tanks arranged in a square within the envelope of the Ariane-5 fairing.



page 231 of 317

sValue Unit

Propellant MMH/NTOMEOP 2.5 MPaTank Material Ti-6Al-4VVolume 2.06 m3

Diameter 1.56 mMinimum Wall Thickness 2.90 mmTank Mass 100.6 kgQuantity 4 Table 7-46: CTV propellant tank characteristics

As regards pressurant tanks, the 22-bar storage pressure is maintained through a helium pressurant system. The helium pressurant is stored in two Composite Over-wrapped Pressure Vessels (COPVs). Table 7-47 shows the pressurant tank characteristics for the CTV UPS:

Value UnitPropellant HeliumMEOP 31 MPaTank Material COPVVolume 0.361 m3

Diameter 0.88 mMinimum Wall Thickness - mmTank Mass 52.0 kgQuantity 2

Table 7-47: CTV pressurant tank characteristics

As regards reaction control system, the RCS uses 16 European Apogee Motors (EAMs), using MMH and NTO storable propellants, which are pressure-fed into each thrusters’ combustion chambers. The engine characteristics are shown in Table 7-25. The RCTs are part of the UPS, and are therefore fed via the same propellant tanks as the main engine. As regards mass budget, Table 7-48 shows the ratio of dry mass to wet mass of the CTV UPS. Here the dry mass includes the pressurant mass, as well as a 10% subsystem margin.

Value UnitDry Mass 1120 kgPropellant Mass 9480 kgTotal Wet Mass 10600 kg

Table 7-48: CTV unified propulsion system mass budget


As regards unified propulsion system, Table 7-49 shows the equipment summary and mass budget, which will be externally mounted to the Crew Transfer Vehicle. The mass budget includes a 10% subsystem dry mass margin and provides the pressurant mass separately.



page 232 of 317

s

Component Mass/Unit (kg) QtyMain Engine (Translunar Injection Engine) 111 1Line Filter 0.23 8Latching Valves (mono) 0.6 4(bi-prop) 0.5 2Pipe work 25 1Pressure Transducer 0.4 7Pressure Regulator 1.5 2(Pyro) Normally Closed Valves 0.35 6(Pyro) Normally Open Valves 0.35 4Reaction Control Thrusters 5 16Isolation Valves 0.5 4Fill/Drain Valve/TP 0.1 14Non-return Valve 0.2 4Propellant tank (ox shell) 100 2Propellant tank (fu shell) 101 2Pressurant tank 52 2Flight Control Orifice (fuel and oxidiser) 0.1 4Thrust Frame Assembly 80 1Adapter 100 1GAM 70 1Dry Mass Margin 10 %

Propulsion System Dry Mass = 1089Propellant Mass = 9480

Pressurant Mass = 31Propulsion System Wet Mass = 10600

Table 7-49: LOI propulsion system equipment summary and mass budget

7.4.5 Conclusions

The design has been based on existing technology with minor modifications. Many propulsion options have been considered for each of the different elements. A design of the AOCS system for the Hub has been proposed. More analysis in this area should be carried out in the future, as no refuelling system has been analysed in the present study. It is assumed that commonality is a main design driver for the transfer stage propulsion systems as there are many in number.

7.5 Hub - AOCS 7.5.1 Introduction

The orbit of the Hub is LLO circular with an altitude of 100 km. In contrast with the International Space Station (ISS), the Hub will operate without a crew for extended periods of



page 233 of 317

stime. Also, the ratio of the mass properties between the Hub and the vehicles docking with it is lower than that of the ISS and the Space Shuttle, the Progress resupply vehicle or the ATV. The design of the AOCS system of the Hub therefore presents some unusual challenges. The low Earth orbit (LEO) assembly of the Hub and its components is not considered here since it is the same as that of the LEV. Only the LLO part of the Hub mission is treated here.


The requirements used for the design of the Hub AOCS system for the LLO part of the mission are that:

• After acquisition of LLO, the Hub shall be given an initial angular rate compatible with the mission profile.

• The Hub’s AOCS shall have the capability to compensate the perturbing forces and torques in LLO.

• The Hub’s AOCS shall handle the large excursion of the CoM and large variations in the MoI during the entire mission.

• The Hub’s AOCS shall be able to command switching to a safe mode compatible with the crew survivability requirements.

The design drivers for the Hub AOCS system identified during this study are:

• The GNC should allow an operator to take over the control of the S/C or at least allow the operator to switch the S/C to a safe mode. (The operator can be the crew on board, or remotely located.)

• To the largest extent possible, the Hub’s GNC HW and SW should be common to that of the LEV.

• The sensors of the Hub should be highly reliable and should be easily integrated in a system that operates autonomously for extended periods of time.

Note that the above requirements and design drivers are not exhaustive. They should be treated as a set of guidelines for future studies rather than an unchangeable set of requirements.


It has been assumed that all Hub orbit maintenance manoeuvres and perturbation compensations are performed autonomously. It has been assumed that only thrusters are used as actuators of the Hub AOCS system.


The angular rate of the orbit is n = 8.886E-4 rad/s. (The orbital period is approximately two hours.) The same angular rate should be imparted to the Hub after acquisition of the LLO by the AOCS system. This rate should then be maintained during the entire mission. An estimate of the ∆V needed to maintain this angular rate by rejecting when subjected to the solar radiation pressure only has been performed. The Hub mass properties are shown in Table 7-50.



page 234 of 317

s

Mass (kg) 50 000 MoI (kgm2) X Y Z

X 14 180 338 7802 Y 338 10 975 149 Z 865 149 21 293

Table 7-50 Hub mass properties

Assuming that the solar radiation pressure (SRP) flux is 1348 W/m2, that the cross-sectional area of the Hub is 70 m2, that the moment arm of the SRP force is 1.5 m, that the thrusters’ arm is 7.5 m, the required ∆V is between 50 and 100 m/s per year, depending on the orientation. The reflectivity factor has been chosen to be 0.5. Note that a careful analysis of the dynamics of the Hub during the entire mission shall be conducted as early as possible, taking into account the final masses and the evolution of the configuration with time, as different vehicles are docking to the Hub at different times. The ISS has been found to have only 24 stable configurations out of the 112 configurations it would have been through during the five years of in-orbit assembly (see RD[36]). For the LLO part of the mission, a sensor suite of Sun sensors, star trackers, and an IMU platform is proposed.

7.5.5 Equipment list

The Hub’s AOCS equipment list is shown in Table 7-51:

Element 1 Unit Name


1 3 axis IMU (gyro + acc) 1 5.000 Fully developed 5 5.32 Autonomous startracker 2 2.000 Fully developed 5 4.23 Coarse Sun sensor 6 0.100 Fully developed 5 0.64 RVS 1 15.000 Fully developed 5 15.85 DGPS 1 5.000 Fully developed 5 5.3- 0.0 To be developed 20 0.0

5 29.6 5.0 31.1Click on button below to insert new unit


margin

Maturity LevelMASS [kg]Element 1: Habitation Module



Unit Quantity

Table 7-51: Hub’s GNC equipment list

7.6 Hub - structures 7.6.1 Requirements and design drivers

For the structure of the Hub, the following set of requirements and design drivers were taken into account:

• Structural design shall aim for simple load paths, maximise the use of conventional materials, simplify interfaces and easy integration.

• The structure shall be designed to meet the requirements for stiffness under the specified load and boundary conditions, this is to avoid dynamic coupling between the low-frequency launcher and Hub modes.



page 235 of 317

s• The structure shall be of adequate strength to withstand the design loads (ground and test

loads, launch loads and in-orbit loads) without yielding, failing or exhibiting excessive deformations that can endanger the mission objectives.

• The structure shall be protected from micrometeoroid and debris impact to prevent the risk of catastrophic failures.

7.6.2 Assumptions

As regards the stiffness requirement, the following is assumed: the fundamental frequency for each Hub Module, hard mounted at the launcher separation plane, is greater than 9 Hz in lateral direction and greater than 27 Hz in longitudinal direction. As regards strength, compression loads from axial and moment loads during launch and stresses from internal pressure in orbit are assumed to be the dimensioning load-cases for the Hub structure. During launch, an ultimate FS of 1.4 is assumed and 1.1 at yield. For the internal pressure, a higher ultimate FS of 2.0 is applied. The highest steady state acceleration is assumed to be 4.55-g in longitudinal direction and 0.25-g in lateral direction.


The design drivers for the configuration of the Hub are the required pressurised volume (habitability requirement) and the constraint of the available volume inside the launcher fairing. Assuming current Ariane-5 fairing even two modules are not enough to comply with the habitability requirement of 450 m3 volume. Therefore it was decided to choose beside a rigid HM, an inflatable HM to increase the available volume in orbit. For the structure of the rigid HM and for the backbone structure of the inflatable HM aluminium stringer/skin with frames is chosen. As a first estimation for the mass of the inflatable HM skin, a foam-rigidised structure is chosen, see Figure 7-19 (from ref. RD[41]). The flexible mesh core material is impregnated with a gelatin resin between membranes of a sealed structure during the fabrication process. The system remains pliable during the stowed configuration. When the structure is deployed and the wall cavity is vented to vacuum, the gelatin resin’s moisture escapes causing foam to expand and harden the mesh core. A micrometeorite shielding outer layer is added, along with an inner, aluminium flame barrier. The mass budget of the inflatable skin is shown in Table 7-52.

Figure 7-19: Layout inflatable skin



page 236 of 317

sSkin Item Mass/Area [kg/m2]

Outer cover and thermal coating 0.332 Four ply fibreglass 0.693 Three adhesive interlayers 0.146 Mesh core and gelatin resin 3.662 Inner thermal coating 0.195 0.1 mm aluminium flame barrier 0.28 MDPS (from ref. RD[4]) 3.012 Total 8.321

Table 7-52: Mass budget of skin inflatable structure

7.6.4 Radiation shielding

Depending on the duration and type of mission, the radiation dose limits that an astronaut can receive during the mission are different. These requirements are translated in terms of the ratio mass per area (g/cm2) of the protection material. For this mission, the requirement in terms of shielding was 9 g/cm2 for the whole spacecraft and 20 g/cm2 for the stormshelter. To determine the mass that needed to be added to fulfil the shielding requirements, the amount of protection provided by the mass already available on board was estimated. The mass available on board to provide shielding protection is shown in Table 7-53:

Material Mass [kg] Hub Skin + Stiffening 1989 Internal Equipment 14628 Hub Debris Shielding 456 MLI 96 Water 2257 Total mass 19426

Table 7-53: Mass budget for radiation protection

Shielding effectiveness depends largely upon the conductivity of the material. Materials with low Z are considered to be effective. The density ratio to convert from g/cm2 to material shielding thickness was used for preliminary calculations, that is, 1 cm of H2O is generally equivalent to a 1 cm thick slab of water, or 4 mm of aluminium. The mass required for a certain shielding requirement is therefore independent of the type of material; only the thickness will be different from one material to another. Two different configurations for the shielding protection provided by the mass available on-board were analysed. The shielding effectiveness of each configuration, for nominal protection and for the stormshelter is shown in Table 7-54.



page 237 of 317

sRigid Shell

Skirt - H2O2257 kg

MLI

Structural Skin additional 2194 kg

Debris Shielding Closures - H2O

14975 kg 2257 kg additional 2194 kg

Skirt Skirt 19.23 g/cm2 19.5 g/cm2

Closures Closures 22.33 g/cm2 21.98 g/cm2

Closures - additional 15%

Int. Equip.

Skirt 180.2 kg

Skirt 117.4 kg

Storm Shelter

20 g/cm2

Skirt - additional 15% Int. Equip.

Protection provided 9.9 g/cm2

Mass to be added ------------------

Shielding Requir. 9 g/cm2

Int. Equip. (68% of total Equip.)

Mass available

Table 7-54: Level of protection provided by the mass on board

Given the assumption that 68% of the total internal equipment is located in the rigid shell of the Hub, and taking into account the protection provided by the MLI, the structural skin and the debris shielding, a total of 14 975 kg are available to use as shielding. This represents a protection of 9.9 g/cm2 for the rigid shell of the Hub, so the requirement is achieved and no extra mass needs to be added to this part of the Hub. As regards the two possibilities analysed for the stormshelter, the one consisting of enlarging the protection on the skirt by means of the 2257 kg of water available and locating 15% of the internal equipment on its closures, is the most attractive. With this configuration, 117.4 kg of extra mass need to be added to the skirt of the stormshelter to guarantee the required protection of 20 g/cm2. The consumables onboard were not taken into account in these calculations because it is not yet known which percentage can be used as shielding protection. Note that the shielding effectiveness of vehicle skin or equipment cases with metal walls of any reasonable thickness is limited by the apertures, joints and others discontinuities, rather than the metal itself. Radiation shielding for the inflatable module has not been analysed in the present study.

7.6.5 Budget

The structure mass budget of the Hub is shown in Table 7-55.



page 238 of 317

s

[kg] [%] [kg]

Habitation Module_skin Rig Mod 1 2226 10 2449Stiffening Rig Mod 1 1113 10 1224Storm Shelter Added Mass Rig Mod 1 117 10 129MDPS Rig Mod 1 456 10 501Attachment tanks/equipment etc Rig Mod 1 200 10 220Ring Frames Rig Mod 1 350 10 385Interior and racks Rig Mod 1 350 10 385Adapter Rig Mod 1 263 10 290Inflatable skin Infl Mod 1 1549 20 1859Flame barrier skin Infl Mod 1 86 10 95MDPS Infl Mod 1 991 10 1090Skin Backbone Cylinder Infl Mod 1 1083 10 1191Stiffeners Backbone Cylinder Infl Mod 1 542 10 596Ring Frames Backbone Cylinder Infl Mod 1 180 10 198Interior and Racks Infl Mod 1 350 10 385Attachment Tanks, equipment etc. Infl Mod 1 200 10 220Adapter Infl Mod 1 107 10 118Airlock Infl Mod 1 300 10 330

10463 11 11665


Table 7-55: Mass budget of habitation modules

7.7 Hub - communications 7.7.1 Requirements and design drivers

In addition to the requirements listed in Chapter 4, Mission Architectures, there are Hub data requirements, as shown in Table 7-56:

Uplink[Mbps]

Downlink[Mbps]

Hub <-> Earth 17.5 20.2 Hub <-> EarthContingencies 10 0.4

Table 7-56: Hub data requirements


See section 6.10.2.



page 239 of 317


7.7.3.1 G/S assumptions

In general, the communications subsystem is similar to the LEV. See relevant sections for details on the design.

7.7.3.2 Communications availability

7.7.3.2.1 Communications availability using G/S

The same as for LEV.

7.7.3.2.2 Communications availability for far Moon side

See Chapter 4, Mission Architectures.

7.7.3.3 EVA

For EVAs and communications inside the Hub, OFDM is proposed (see section 6.10.2.7).

7.7.3.4 Architecture and frequency plan

See Chapter 4, Mission Architectures.

7.7.3.5 UHF link

See section 6.10.3.5.

7.7.3.6 Link characteristics

Table 7-57 provides a summary of the Hub links:

Link Ka+ band1

X-band (HGA) UHF2

X-band (LGA) Uplink Downlink Uplink Downlink Uplink Downlink Uplink Downlink

Frequency 40-40.5 GHz 37.5-38 GHz 7.19-7.235 GHz 8.45-8.5 GHz 410-420 MHz 410-420 MHz 7.19-7.235 GHz 8.45-8.5 GHz

Tx power 40W 10W 71W 10W 5W 5W 71W 10W

Modulation NRZ/PSK/PM GMSK. BTb=0.25


OQPSK3


BPSK/PM BPSK/PM SRRC (α: 0.5)

Filtered OQPSK


Coding


+ RS (255, 223)

Turbo Coding ¼ RS (255, 223) RS (255, 223)


+ RS (255, 223)

Concatenated: Convolutional + RS (255, 223)

RS (255, 223) RS (255, 223)

FER Negligible Negligible Negligible Negligible BER=10-5 BER=10-5 Negligible BER=10-6 Bit rate

(worst case) 40 Mbps 100 Mbps 10 Mbps 10 Mbps 256 kbps 256 kbps 10 Mbps 400 Kbps

Table 7-57: Links description

1 Atmospheric attenuation of 4.5 dB, it corresponds with an elevation of 10° and G/S availability of 90%. 2 Max distance Hub-LEV of 368 km, corresponding to a Hub elevation over the horizon of 10°. 3No standard modulation exists for uplink high data rates.



page 240 of 317


7.7.4.1 Antenna summary

Table 7-58: Antenna summary

7.7.4.2 Equipment summary

Element 1 Unit NameClick on button below to insert

new unit1 UHF omni antenna 3 1.2 Fully developed 5 3.82 Ka+ / X-band dish antenna 2 2.5 To be modified 10 5.53 UHF transceiver 3 3.0 Fully developed 5 9.54 Ka+ / X-band transponder 3 4.6 To be developed 20 16.65 SSPA Ka-band 3 0.8 To be modified 10 2.66 SSPA X-band 3 0.8 Fully developed 5 2.57 Global RFDU unit 1 5.0 To be developed 20 6.08 X-band omni antenna 3 0.4 Fully developed 5 1.39 EVA 802.11g router/rxer 4 0.3 To be developed 20 1.410 EVA 802.11g antenna 4 0.1 To be developed 20 0.511 Ka-band transponder 2 4.0 Fully developed 5 8.412 Ka-band dish antenna 1 2.5 Fully developed 5 2.613 Ka-band RFDU 1 1.0 To be modified 10 1.1- 0.0 To be developed 20 0.0



margin

Maturity LevelMASS [kg]Element 1: Habitation Module



Unit Quantity

Table 7-59: Hub equipment summary

7.8 Hub - data handling This section describes the basic requirements, design drivers and baseline design description for the Data Handling System (DHS) of the Hub module. Commonality with the LEV design has been sought whenever possible.


See section 6.11.1.

7.8.2 Main issues and proposed building blocks

See section 6.11.2.


A fully decentralised data handling system is proposed. The approach for the architecture is very similar to the LEV. The network characteristics are identical to the LEV so that when both spacecrafts are docked, a connection is established between both networks allowing for communication and remote control from one to the other. Wireless networks are also proposed for internal use within the Hub for facilitating the interface of the crew with the several on-board computers.

Element Type of antenna Band Pointing device Transmitted

power (W)

Hub

Two dish 50 cm One dish 40 cm for TDRSS Three LGA antennas (quadrifilar helix) Three omni-directional UHF antennas

Ka+ and X band Ka band X-band UHF

2 DOF Steering 2 DOF Steering None None

10 10 10 5



page 241 of 317

sConsidering the long duration of the Hub mission and for increased reliability, spare/redundant units have been included in the design for the General System Computer, the Housekeeping Unit and the Safeguard Data Recorder. Figure 7-20 shows the data handling system architecture:

Habitation Module

GN&C Computer Safeguard Data Recorder UNITS 2.00dim 266.00 190.00 168.00 mm3 dim 266.00 190.00 126.00 mm3mass 6.60 kg mass 4.90 kgpow 30.00 W pow 5.00 W

DM on board bus

General System Computer UNITS 2.00 Mass Memory Unitdim 266.00 190.00 273.00 mm3 dim 266.00 190.00 126.00 mm3mass 10.85 kg mass 4.90 kgpow 34.89 W pow 18.33 W

Analog I/OHousekeeping Unit UNITS 2.00dim 266.00 190.00 210.00 mm3 Digital I/Omass 8.30 kgpow 6.02 W Serial Lines

Communication Unitdim Included in General System Comp Transponders --> Communications Subsystemmass But with an independent BUS I/Fpow

MMI+SCREENDistributed control bus Computerspacecraft backbone keyboard

High Speed Data Network MMI+SCREENCommand & Control Network Computer

keyboard

MMI+SCREENMicro RTUs UNITS 15.00 Computerdim 133.00 95.00 42.00 mm3 keyboardmass 6.00 kgpow 12.00 W

MMI+SCREENComputerWireless I/Fkeyboard

Wireless Access pointdim - - - mm3 EVAmass - kg Computerpow - W Wireless I/F

HANDHELDDeviceWireless I/F

Backbone network connections in docking ports HANDHELDDevice

Backbone network connections to LEV Wireless I/F

Figure 7-20: Hub habitation module’s DHS architecture



page 242 of 317



unit1 General System Computer 2 10.9 To be developed 20 26.02 GN&C Computer 1 6.6 To be developed 20 7.93 Safeguard Data Recorder 2 4.9 To be developed 20 11.84 Mass Memory Unit 1 4.9 To be developed 20 5.95 Micro RTUs 1 6.0 To be developed 20 7.26 Man Machine Interface 5 5.3 To be developed 20 31.77 Housekeeping Unit 2 8.3 To be developed 20 19.98 Cameras 4HQ + 20LQ 1 1.8 To be developed 20 2.29 EVA Computer & Handheld devices 4 1.0 To be developed 20 4.8- 0.0 To be modified 10 0.0


Unit QuantityElement 1: Habitation Module



margin

Maturity LevelMASS [kg]


Table 7-60: Data handling equipment list for the Hub

7.9 Hub - mechanisms 7.9.1 Requirements and design drivers

As a result of the Hub’s configuration, the following necessary mechanisms and their requirements can be derived:

• Habitation Module (Hub): o Crew Ingress/Egress Hatches at Docking ports o Hatch diameter 0.9 m o Four docking ports o Inter-module connection:

− EVA Egress Hatches - Internal and External • Hatch Diameter 1.2 m.

o Power Generation System: − Deployable Solar Arrays − Panel Hold-down & Deployment:

• Required Surface Area- two wings of 61 m2 • Total Panel Mass 241.8 kg incl. 20% margin

− Solar Array Wing Rotation: • 360° continuous rotation – Sun tracking

− Potential for wing restowage and latching capability- this is dependent upon the loading introduced by propulsive manoeuvres

o Thermal System: − Deployable Radiator Panels/Arrays − Panel Hold-down & Deployment:

• Required Surface Area- two wings of 77 m2 • Total Panel/Structural mass 847 kg incl. 10% margin

− Radiator Panel Rotation: • 360° continuous rotation – Sun tracking

− Potential for radiator panel restowage and latching capability- this is dependent upon the loading introduced by propulsive manoeuvres.



page 243 of 317

so Berthing and release capability:

− Inter-stage berthing for Hub Propulsion Module I/F. − Propulsion Module Stack I/F- not considered in this study

o Berthing and docking capability: − Berthing & Docking/Undocking in Lunar Orbit- AV to Hub. − Berthing & Docking/Undocking in Lunar Orbit- Resupply Vehicles. − Berthing & Docking/Undocking in Lunar Orbit- Crew Transfer.

o Communication system: − Antenna Pointing and Tracking Mechanism:

• Type 1: Antenna diameter: 0.4 m Antenna mass: 2.5 kg estimated Coverage: 180° hemispherical. Pointing accuracy: 10° Boom mounted at 0.5 m Number of antennas- 1

• Type 2: Antenna diameter: 0.5 m Antenna mass: 3 kg estimated Coverage: 180° hemispherical. Pointing accuracy: 2° Boom mounted at 0.5 m Number of antennas- 2


7.9.2.1 Power Generation System

The following assumptions have been taken: • The number of arrays shall be minimised • The mass shall be minimised • The technology chosen shall ensure that stowage of the array is possible

The following solar array deployment systems are available:

1. Advanced Rigid Arrays (ARA) 2. Polar Platform arrays.

The key features of the two concepts are:

1. Advanced Rigid Arrays: • Typically four or five panel wings with surface area of 30-35 m2 (typical panel size

2.5 m x 2.75 m). • Spring driven, single direction deployment, with latched panels for in-flight wing

stiffness. • No re-stowed latching capability

2. Polar Platform Arrays: • Up to 16 panel capability with surface area of up to 80 m2 (typical panel size 5 m x 1

m, current qualification status of 14 panels per wing).



page 244 of 317

s• Motorised, cable actuated deployment with re-stowage capability (not yet qualified). • No restowed latching capability (yet).

The main characteristics indicate that the polar platform-type array would better suit this application when considering the assumptions made above (2+ array wings against 5+ for the ARA type considering a maximum size wing). Additionally, the requirement for restowage is facilitated by a motorised deployment system and the PPF array can be easily modified for this capability.

7.9.2.2 Communications system

The following assumptions have been derived as a result of the study: • All boom-mounted antennas require tracking capability. • Tracking can be realised with two perpendicular rotational axes.

No trade-off has been performed. The choice of the chosen mechanism has been made based upon the available systems and the requirements stated earlier.

7.9.2.3 Vehicle connections

Two systems have been considered: 1. Russian Docking System (see left part of Figure 7-21). 2. International Berthing and Docking Mechanism (see right part of Figure 7-21):

Figure 7-21: Russian docking system, International berthing and docking mechanism

The two systems can be briefly characterised as follows: 1. Russian Docking System:

• Both halves of the system are integrated in the hatch doors • “Male” half situated on approaching vehicle • Receptacle situated on the Hub • Internal redundancy of systems • System-level redundancy not complete. Redundant receptacle can be provided, but no

redundant probe can be mounted on approaching vehicle

2. International Berthing and Docking Mechanism (IBDM): • Androgynous system- identical mechanism mounted to both vehicles • Full redundancy of system provided



page 245 of 317

s• Full internal mechanism redundancy • Triple redundancy for release/emergency release • Mechanism independent of hatch door • Hatch door diameter currently limited to ingress/egress suitability (∅813 mm) but an

upgrade is in development to increase the door diameter to 1.0-1.2 m


7.9.3.1 Crew ingress/egress hatches

Sealable hatches are required for the following I/Fs: • All docking ports • Inter-module connection- one on either module • Internal and external airlock hatch.

A total number of eight hatches is required. The hatch diameter is sized to about ∅0.9 m for all hatches through an IBDM. The airlock hatches are sized to about ∅1.2 m to allow a fully suited astronaut to pass through. The hatch will require latch and seal mechanisms. Mass estimates is realised using a ‘simple geometry’ model and material mass properties.

7.9.3.2 Berthing capability

For in-orbit assembly of the Propulsion Module to the Hub modules, the Common Berthing Mechanism shall be the baseline adapted to suit the purpose of a non-pressurised inter-vehicle connection.

Figure 7-22: Common berthing mechanism

7.9.3.3 Vehicle docking

The International Berthing and Docking Mechanism (IBDM) shall be implemented all for the docking interfaces on the Hub. The IBDM in its current form has the following mechanical characteristics based upon design analysis:

• Outside diameter: 1.371 m; Inside diameter: 0.813 m (effective pass through diameter); H 0.254 m (retracted).

• Mass estimate: 304 kg.



page 246 of 317

s• Interface loads (at the sealing interface) acting simultaneously while docked (Flight-

limit) o Axial load: 5338 N o Shear load: 4448 N o Bending moment: 9039 Nm o Torsion moment: 7909 N*m

• Internal pressure: 110316.1 Pa • Life: 15 years; functional life: 20 berthing/unberthing or docking/undocking cycles

The advantages of the IBDM system are:

• It is an androgynous system implemented on both sides of the mated interface (no male- female connections).

• The system is present on both sides of the I/F, so a fully redundant system is realised. Currently, further IBDM development is focused upon increasing the available internal diameter to enable racks to be passed through the IBDM. The intended internal diameter will increase to between 1.0 and 1.2 m (from the current 0.81 m) without increasing the structural ring dimensions by placing all the structural connections outside the pressurised volume rather than inside. To calculate mass, the ratio of the internal diameter increase has been used. The changes to the IBDM are not major structural changes, so 50% of the calculated increase has been added to the original mass to give a mass estimate for the increased size IBDM of 380 kg. As there are five IDBMs being used on the system and commonality is whished in the system, the increased size IBDM has not been implemented due to the resulting mass penalty/increase. The IBDM function is supported by eight electronic boxes of size about 0.4 m x 0.25 m x 0.25 m with a mass of 8 kg each. The loading on the IDBM has been calculated for the types of vehicle interfaced to the Hub and also during all phases of the mission. The results are shown in Table 7-61: ATV

Docked System (Kg) Total (Kg)15562.5 15562.5 Assumed 75% of Launch Mass

CoM position 4.897 m Assumed 50% of Length ConversionAllowable Shear 4448.2216 N 1000 lb.f 4.4482216Allowable Bending 9038.8 Nm 80000 lb.f-inch 0.112985

Mass. Vehicle (Kg) Thrust (N) Acceleration (m/s)Force applied at CoG/Shear Load

IBDM Bending Moment

Margin of Safety Bending

TLI. 1st Start 87995 64800 0.736405478 9573.271209 19146.54242 -0.527914764TLI. 1st End 65682 64800 0.986571663 12825.43163 25650.86325 -0.647621996TLI. 2nd Start 61205 64800 1.058737031 13763.58141 27527.16281 -0.671640697TLI. 2nd End 39667 64800 1.633599718 21236.79633 42473.59266 -0.787190124LOI. 1st Start 34600 64800 1.87283237 24346.82081 48693.64162 -0.814374122LOI 1st End 28517 64800 2.272328786 29540.27422 59080.54844 -0.847008868

Station Keeping- basic Hub + ATV 69562.5 800 0.011500449 149.5058401 299.0116801 29.22891947Station Keeping - Max. Mass co 170000 800 0.004705882 61.17647059 122.3529412 72.87480769



page 247 of 317

sKlipper/CTV

Docked System (Kg) Total (Kg)13000 13000

CoM position 2 m Estimate ConversionAllowable Shear 4448.2216 N 1000 lb.f 4.4482216Allowable Bending 9038.8 Nm 80000 lb.f-inch 0.112985


IBDM Bending Moment



Station Keeping- basic Hub 54000 800 0.014814815 192.5925926 385.1851852 22.46611538Station Keeping incl. LEV 170000 800 0.004705882 61.17647059 122.3529412 72.87480769 LEV

Docked System (Kg) Total (Kg)50000 50000

CoM position 6 m ConversionAllowable Shear 4448.2216 N 1000 lb.f 4.4482216Allowable Bending 9038.8 Nm 80000 lb.f-inch 0.112985


IBDM Bending Moment


Station Keeping- basic Hub 54000 800 0.014814815 192.5925926 385.1851852 22.46611538Station Keeping incl. LEV 170000 800 0.004705882 61.17647059 122.3529412 72.87480769

Table 7-61: IBDM load cases and results

At no time during the lunar transfer shall a vehicle be docked to the Hub. For all stationkeeping manoeuvres, the IBDMs are not overloaded.

7.9.3.4 Communication antennas

Current APM systems are able to meet the pointing requirements for the antenna. A suitable unit has been chosen and an estimate of the deployable boom mass has been made. A schematic of a typical boom mounted deployment and pointing mechanism for the TV antennas is shown in Figure 7-23:

Deployment Hinge

Axial Hinge

Azimuth Hinge

Antenna Mounting I/F

Figure 7-23: Antenna pointing mechanism

For a complete 180° hemispherical coverage, the axial and azimuth axes shall rotate through 180°. An analysis of the loading applied to the antenna boom gives the results shown in Table 7-62:



page 248 of 317

s HGA-1 HGA-2Mass- Dish 3 Kg Mass 2.5 KgMass APM 9.4 Kg Mass APM 9.4 KgTotal Mass 12.4 Kg Total Mass 11.9 KgDia 0.5 m Dia 0.5Boom Length 0.5 m Boom Length 0.5 Kg

Mass. Vehicle (Kg) Thrust (N)

Acceleration (m/s)





TLI. 1st Start 87995 64800 0.7364 9.1314 4.5657 8.7632 4.3816TLI. 1st End 65682 64800 0.9866 12.2335 6.1167 11.7402 5.8701TLI. 2nd Start 61205 64800 1.0587 13.1283 6.5642 12.5990 6.2995TLI. 2nd End 39667 64800 1.6336 20.2566 10.1283 19.4398 9.7199LOI. 1st Start 34600 64800 1.8728 23.2231 11.6116 22.2867 11.1434LOI 1st End 28517 64800 2.2723 28.1769 14.0884 27.0407 13.5204

Station Keeping- basic Hub 54000 800 0.01481 0.1837 0.0919 0.1763 0.0881Station Keeping incl. LEV 170000 800 0.00471 0.0584 0.0292 0.0560 0.0280

Table 7-62: Antenna boom loads analysis

The results show that the antenna root hinge will be loaded to about 14 Nm. A deployed antenna system should be able to survive this level of loading so the antennas can remain deployed throughout the mission.

7.9.3.5 Power Generation System

The baseline system shall be the polar platform type of solar array. The characteristics of the array are as follows:

• Individual panel size 5 m x 1 m • 13 active panels per wing; a wing requires 14 panels for deployment system to function

correctly • 12 panel hold-downs per wing; two deployment beam hold-downs per wing. • Total area 65 to 70 m2 per wing

The required improvements to be made to the array are as follows:

• Inclusion of the retraction/re-stowage capability; function already available, but not qualified.

• Inclusion of a relatching and release capability in the stowed configuration. Each solar array wing shall be mounted via a root hinge to a solar array drive mechanism, such as the SEPTA 31 used for the PPF Solar Array on Envisat. This model has been selected but is likely to be overqualified for the application. An analysis of the loading on the deployed solar array gives the results shown in Table 7-63:



page 249 of 317

sSolar Array Wing Mass

Deployment System (Kg) Panel/Cells (Kg) Total (Kg)31.427 241.8 273.227

CoM position 7.5 mAllowable Shear 24.66666667 N (Derived from Bending allowable, not actual shear)Allowable Bending 185 Nm

Mass. Vehicle (Kg) Thrust (N) Acceleration (m/s)SA Force applied at

CoG/Shear LoadSA Bending


Bending



Table 7-63: Solar arrays load analysis

The results show that the array cannot remain deployed during lunar transfer propulsive events. During lunar orbit stationkeeping manoeuvres, the array can remain deployed.

7.9.3.6 Thermal system

The radiator panels shall be deployed using a deployment system similar to that of the solar wings. Additionally, a similar SADM will be employed. The characteristics of the radiator are as follows:

• Individual panel size 5.2 m x 1 m • 15 active panels per wing; wing requires 16 panels for deployment system to function

correctly • Initial estimate:12 panel hold-downs per wing; two deployment beam hold-downs per

wing. • Total area 78 to 83 m2 per wing

The required improvements to be made to the array are as follows:

• Increase the structural capability of the deployment system due to the significantly heavier radiator panels.

• Inclusion of the retraction/restowage capability; function already available, but not qualified.

• Inclusion of a relatching and release capability in the stowed configuration. A mass estimate of the deployment system has been realised based upon the PPF SA mass estimates and applying a scale factor calculated from the difference of the SA structural/panel mass and the radiator structural mass. The scale factor is 3.52. An analysis of the loading on the radiator array gives the results shown in Table 7-63:



page 250 of 317

sRadiator Array Wing Mass

Deployment System (Kg) Panel (Kg) Total (Kg)137.52 847 984.52

CoM position 8.5 mAllowable Shear 25 N (Derived from Bending allowable, not actual shear)Allowable Bending 185 Nm

Mass. Vehicle (Kg) Thrust (N) Acceleration (m/s)SA Force applied at

CoG/Shear LoadSA Bending


Bending



Table 7-64: Radiators load analysis

Note that the analysis has been performed assuming the current structural capability of the PPF SA. If the capability of the deployment system is improved, the margins shown will increase. The results show that the array cannot remain deployed during Lunar Transfer propulsive events (even with improved capability, there is a factor of 100 difference). During lunar orbit stationkeeping manoeuvres, the array can remain deployed but with a small margin for the minimum Hub mass case.

7.9.4 Budgets

Element 1 Unit Name


1 Hatch Door-Egress External 6 16.0 20 115.2 0.9 0.0102 Hatch Door Locking Mechanisms- Egress External 6 75.0 20 540.0 0.90 0.8 0.053 Docking Mechanism- IBDM 4 334.4 10 1471.4 1.810 1.200 0.2544 Electronic Box- IBDM 32 8.8 10 309.8 0.40 0.25 0.255 Common Berthing Mechanism- Active 1 311.0 20 373.2 2.0 1.8 0.1906 Common Berthing Mechanism- Passive 1 177.0 20 212.4 2.0 1.8 0.3437 Electronic Box- Common Berthing Mech 1 8.0 20 9.6 0.4 0.25 0.258 Solar Array Hold-down Systems 2 6.5 20 15.69 Radiator Array Hold-down systems 2 6.5 20 15.610 Antenna Pointing Mechanism- APM+ Boom 3 9.4 20 33.711 Electronic Box- APM+ Boom Deploy 3 5.0 20 18.012 Berthing Mechanism- Prop. Active 1 158.9 20 190.7 2.8 2.6 0.213 Berthing Mechanism- Prop. Passive 1 143.0 20 171.6 2.8 2.6 0.314 Electronic Box- Prop. Berthing Mech 1 8.0 20 9.615 Radiator Array Deployment Mechanism- SDM 2 129.3 20 310.316 Radiator Array SDM/Panel Hinges 2 17.9 20 43.117 Radiator Array Yoke Panel 2 14.1 20 33.918 Radiator array Root Hinge 2 17.6 20 42.419 Solar Array Deployment Mechanism- SDM 2 36.7 10 80.720 Solar Array SDM/Panel Hinges 2 4.9 10 10.821 Solar Array Yoke Panel 2 4.4 10 9.7 1.0 2.5022 Solar array Root Hinge 2 5.5 10 12.1 0.5 0.30 0.15023 Solar Array Drive/Rotation Mechanism (SEPTA 31) 2 44.00 10 96.824 Solar Array Stowed Latch(s) 2 4.0 20 9.625 Solar Array Drive Electronics 1 4.7 10 5.226 Radiator stowed Latch(s) 2 8.0 20 19.227 Radiator Drive/Rotation Mechanism (SEPTA 31) 2 44.0 20 105.628 Radiator Drive Electronics 2 4.7 20 11.329 Hatch Door-Egress Airlock 2 35.0 20 84.0 1.25 0.01030 Hatch Door Locking Mechanism-Egress Airlock 2 170.0 20 408.0 1.20 1.35 0.050- 0.0 20 0.0

30 4125.3 15.6 4768.8-Click on button below to insert new unit


margin

DIMENSIONS [m]MASS [kg]Dim3 Height

Dim2 Width or d

Element 1: Habitation ModuleMargin Total Mass

incl. marginDim1

Length or D


Unit Quantity

Table 7-65: Hub mechanisms equipment list and mass budget



page 251 of 317

sElement 1 Unit Name OEM OEM OEM TMIM TMIM TMIM OMM OMM OMM

Click on button below to insert new unitPon Pstby Dc Pon Pstby Dc Pon Pstby Dc

1 Hatch Door-Egress External2 Hatch Door Locking Mechanisms- Egress External3 Docking Mechanism- IBDM 1806.0 0.0 0.0015 1806.0 0.0 0.00154 Electronic Box- IBDM 152.0 0.0 0.00876 152.0 0.0 0.008765 Common Berthing Mechanism- Active 575.0 25.0 0.0066 Common Berthing Mechanism- Passive7 Electronic Box- Common Berthing Mech 25.0 0.0 0.0068 Solar Array Hold-down Systems9 Radiator Array Hold-down systems

10 Antenna Pointing Mechanism- APM+ Boom 14.0 0.0 100.0 14.0 0.0 100.0 14.0 0.0 100.011 Electronic Box- APM+ Boom Deploy 10.0 0.0 100.0 10.0 0.0 100.0 10.0 0.0 100.012 Berthing Mechanism- Prop. Active13 Berthing Mechanism- Prop. Passive14 Electronic Box- Prop. Berthing Mech 25.0 0.015 Radiator Array Deployment Mechanism- SDM 30.0 0.0 0.003 30.0 0.0 0.003 30.0 0.0 0.00316 Radiator Array SDM/Panel Hinges17 Radiator Array Yoke Panel18 Radiator array Root Hinge19 Solar Array Deployment Mechanism- SDM 30.0 0.0 0.003 30.0 0.0 0.5556 30.0 0.0 0.003120 Solar Array SDM/Panel Hinges21 Solar Array Yoke Panel22 Solar array Root Hinge23 Solar Array Drive/Rotation Mechanism (SEPTA 31) 10.0 2.0 100.0 10.0 2.0 100.0 10.0 2.0 100.024 Solar Array Stowed Latch(s)25 Solar Array Drive Electronics 5.0 2.0 100.0 5.0 2.0 100.0 5.0 2.0 100.026 Radiator stowed Latch(s)27 Radiator Drive/Rotation Mechanism (SEPTA 31) 10.0 2.0 100.0 10.0 2.0 100.0 10.0 2.0 100.028 Radiator Drive Electronics 5.0 2.0 100.0 5.0 2.0 100.0 5.0 2.0 100.029 Hatch Door-Egress Airlock30 Hatch Door Locking Mechanism-Egress Airlock-

0.0 2697.0 33.0 114.0 8.0 2072.0 8.0Click on button below to insert new unit

ND POWER SPECIFICATION PER MODE PPEAK AND POWER SPECIFICATION PER Ppeak

Element 1: Habitation Module


Unit

Table 7-66: Hub mechanisms equipment list and power budget

7.10 Hub - power 7.10.1 Requirements

The main requirements of the power subsystem design for the Hub are: • Mission duration of 10 years in LLO orbit • A safety level consistent with a crewed vehicle • A power subsystem that fits with the assembly constraint (volume limitation,

mounting…) • The use of technologies expected to be qualified in 2015 for a mission in 2020-2025

For this study, only technologies already qualified in space (with improvements expected in the coming years) are taken into account. The confidence in this subsystem is higher compared to others possible designs for which qualification has lower probability to be reached in time.

7.10.2 Assumptions

After assembly of the Hub modules with their relevant propulsion modules in LEO, the Hub will be injected into a LLO 100 km circular orbit for about 10 years. During these 10 years, replacement of components may be possible via cargo missions. A 100 km circular polar orbit around the Moon has an orbit duration of 127.6 minutes with eclipses that can extend to 45.2 minutes depending on the orbit selection and the date. In addition to these frequent eclipses encountered, Earth eclipses also occur (around twice per year). These eclipses are longer and their durations vary with the years. As a sizing case, six hours of occultation is taken into account for sizing the power storage of the Hub.



page 252 of 317

s The Hub power system will have to be sized to provide power to the elements that will be docked to it, i.e., two LEVs, the CTV and the cargo vehicles.

7.10.2.1 Mission assumptions

To cover failures such as an attitude loss of the spacecraft, the power subsystem is designed to cope at any moment with a total solar power generation failure of six hours. Due to the tight timeline of this study, a detailed list of modes has not been assessed for the power subsystem. For example, the assembly phase in LEO of the Hub was not considered. However, since the Hub should be unmanned during the assembly phase in LEO and even during the cruise to LLO, this phase is not a sizing case for the power subsystem design. The identified design case is:

• EOL (10 years in LLO) • LEV docked and power supplied by the Hub • CTV docked and power supplied by the Hub • Crew on board the Hub • Longest possible Moon eclipse (45.2 minutes) on the 127.6 minutes orbit • Longest possible Earth occultation (6 hours with margin included)

7.10.2.2 Power assumptions

The power budget for the worst-case mode defined in this section has been assessed on similar previous studies:

• The CTV power consumption is derived from the Soyuz capsule and the previous lunar CTV study (RD[61])

• The LEV power supply in orbit is based on the LEV safe mode And on the design of the Hub subsystems a margin of 20% has been added for mission operations flexibility and for uncertainties of the bus users (see Table 7-67). The average power consumption required is 6.75 kW (8.1 kW with margin). The switching on of all the units at the same time would require 17.7 kW (21.2 kW with margin). This value will clearly never be reached during the mission and therefore shall not be considered in the design of the power subsystem sources. Concerning the harness and the power distribution units, this peak power value has to be considered. The average power consumption with margin is the sizing value for the power source design. To cope with temporary power requirements over the 8.1 kW on the bus, a unit able to provide the extra power always has to be implemented.



page 253 of 317

sCTV Thermal 2 LEVs Comms Propulsion DHS Life Support Mech

linked linked linked linked linked linked linked linkedPpeak 0 W 0 W 0 W #NAME? 0 W 0 W 0 W 0 W

Pon 166 W 6754 W 0 W 387 W 0 W 80 W 6743 W 2707 WPstdby 66 W 95 W 0 W 17 W 0 W 80 W 0 W 35 W

Duty Cycle 33 % 13 % 100 % 100 % 0 % 0 % 52 % 2 %Tref 259200 min Total Wh 427680 Wh 4132987 Wh 0 Wh 1671840 Wh 0 Wh 345600 Wh 15067703 Wh 384804 Wh





P average 99 W 1402 W 1094 W 387 W 0 W 80 W 3488 W 64 W

Safe Mode

Trans-Moon Injection Mode

Orbiting around Earth Mode

Orbiting around Moon Mode

Eclipse Mode :

LLO Power Requirement on bus: Peak 21.2 kW

Average: 8100 W+20% Bus Margin

LLO Power Requirement on bus: Peak 17.7 kW

Average: 6750 W

Energy Supply Sizing

Table 7-67: Power budget assessment

7.10.3 Trade-offs between technologies for the Hub

7.10.3.1 Power generation

From all the non-nuclear primary energy source technologies for space applications, photovoltaic conversion is the most used, qualified and the most promising for future space missions.

7.10.3.1.1 Photovoltaic cells

In this field, research is taking place to increase the performances of the cells (efficiency conversion, mass, cost, reliability, stored volume). This research also includes development of thin film cells that could fit on flexible or inflatable structures. Such a product should be available in 2015. Two options of photovoltaic cells are computed in this study:

• Rigid cells: o 27% efficiency (AM0 (28°C)) reached now with AsGa TJ cells o 30% expected to be reached in 2015 o 3.3 kg/m2 for the total mass of the PVA

• Thin-film cells: o Still in development for space application but Earth applications already exist o In 2015/2020: 15% efficiency expected (AM0 (28°C)) o A promising candidate is the CIGS cells mounted on thin polyimide foil o 0.6 kg/m2 for the mass of the PVS including the frame but excluding the mechanisms



page 254 of 317

s

Table 7-68: Assumptions for the two solar cells options

Table 7-69 shows the main points that have to be considered when selecting rigid SA cells or thin-film cells:

Advantages Disadvantages Low cost expected Lower Efficiency Monolithic integrated structure (connections…) Lack of space experience Higher EOL Power density Lack of manufacturing experience with space qualified materials High Specific power and stowability No diodes yet incorporated in design Flexible flat cabling not yet incorporated

Table 7-69: Advantages/disadvantages of thin film cells compared to rigid PV cells

7.10.3.1.2 Pointing attitude

In LLO, the Hub attitude is only constrained by having its longitudinal axis in the same direction as the velocity vector. Consequently, the Hub has one degree of freedom: the rotation around its longitudinal axis. To have the solar cells almost Sun pointed for a circular polar orbit, another one-axis Sun pointed mechanism is required, perpendicular to the longitudinal direction. Pointing accuracy and inclination of the polar orbit of the Hub towards the Sun direction requires taking into account a worst-case depointment of 15 degrees. The reliability requirements for a human mission also include the loss of a complete photovoltaic wing. Therefore, an additional redundant wing has to be implemented. Taking into account the final area of solar cells needed, the accommodation difficulties and the mechanisms factors, a configuration with two symmetrical and parallel solar wings supplying in total double the nominal power required is chosen. As a second option, a configuration with four solar panels in a cross configuration around the fairing of the Hub is also possible: the total solar cells surface would be identical to 50% smaller wings but four SADMs are mandatory in this design.

7.10.3.1.3 Mounting and deployment structure of thin-film cells

In addition to the development issue for the thin-film cells, the mounting and deployment structure has also to be optimised in terms of mass. Several deployment of flexible structures have already been studied (RD[52] and RD[53]). The most advanced concepts are:

• Use of a pantograph mechanism:



page 255 of 317

so Based on ERS 1-2/SPOT 1-2-3 solar wings o High mass-to-area ratio of mechanical part: 1.5 kg/m2 for a 39 m2 area

• Use of an inflatable boom: o Conceptual structure performed by Alcatel and Ideamach o High mass: 40 kg for 1 kW LEO: optimisation required o Technology enhancements already expected for years

• Use of a deployable boom: o Based on a telescopic boom o TRW has studied this concept especially for interplanetary spacecraft using electrical

propulsion. The wing was 57.8 m2 deployed for a stowed volume of only 38 dm³ o No European space heritage available, further development studies are required

• Use of a rigid telescopic mast: o A lot of fixation points leading to a stiffness problem for long wings

• Astromast (See also RD[54]): o Space qualified with ERS, Olympus… o ERS wing: 3 m x 20 m long with a mast mass of around 40 kg (0.6 kg/m2 only

without stiffness units) o No problem of stiffness or stability o Accurate sizing of the wings (optimal length and width) are linked to the stiffness

capabilities of the mounted thin films and therefore cannot be computed yet.

Figure 7-24: Deployment and mounting mechanisms for thin-film cells



page 256 of 317

s

Figure 7-25: Astromast illustration with its canister

From all these existing deployment mechanisms, the Astromast is selected according to ESA’s heritage, mass, stiffness and stowed volume. The main disadvantage of the Astromast is the canister height that requires special accommodations. A conservative value of 0.8 kg/m2 is assumed in the solar wing design including also the stiffness elements.

7.10.3.2 Power storage

The mission of the Hub as defined includes many eclipses from the Moon itself or from Earth. The three best identified candidates for the Hub are:

1. Secondary batteries 2. Regenerative fuel cells 3. Flywheels

Note that in addition to the supply of power during the eclipses and manoeuvre phases, the power storage module is also the only reliable power source during a S/C attitude failure.

7.10.3.2.1 Secondary batteries

The most efficient units are the Li-Ion cells with roundtrip efficiency around 94%. The specific energy nowadays is around 100 Wh/kg and this technology is space qualified. An enhancement to 150 Wh/kg is expected by 2015 (RD[64], RD[65] and RD[66]). Depending on the LLO orbit inclination, the battery will need to be sized for 40 000 cycles. About 50% of the nominal capacity should only remain in the cells at EOL. Table 7-70 shows the assessment of the mass budget of the different units of the power subsystem (for 6-hour Earth occultation and 6-hour safe mode, rigid SA and secondary batteries architecture): 1515 kg for the mass of the battery,



page 257 of 317

s403 kg for the rigid solar panels, 135 kg for the electronic modules giving a total mass of 2463 kg with the unit margin included.

Table 7-70: Power subsystem sizing for secondary batteries architecture

7.10.3.2.2 Regenerative fuel cells

These are fuel cells that can also be used in a reversible way: when power is supplied, an electrolysis process takes place and the fuels (H2, O2) are produced. Compared to Li-Ion batteries, regenerative fuel cells have low efficiency (40-60%), which increases directly the size required for the PVAs. Fuel cells already have a space background (Apollo, Space Shuttle) but not the regenerative process. As for the fuel cells, the interest of the regenerative fuel cells for space applications are related to the storage capabilities of the fuels (the types of fuel storage of the regenerative fuel cells are identical to the primary fuel cells). Compared to terrestrial application, the oxygen also needs to be carried. A cryogenic storage of the hydrogen and the oxygen is selected for volume optimisation and safety of the crew. Compared to a gaseous storage of the fuels, a cryogenic storage requires additional units: H2/O2 Driers, a H2 liquefaction unit and an O2 liquefaction unit with associated radiators. 40% is expected for the round trip efficiency of this storage process. The advantages of this solution are mainly the lightweight tanks used with lower levels of pressure requirements: cryogenic storage tanks consists of a spherical aluminium inner pressure vessel and a concentric aluminium outer shell with several dozen layers of MLI and vapour-cooled shields placed between the inner and the outer spheres. To limit the power required by the cooling system dedicated to the hydrogen tanks and also the boil-off, one solution is to keep the maximum fuel in the water liquid state and to perform the electrolysis just prior to an Earth occultation. In the event of needing to refill the fuels, this can also be done easily by simply bringing water and performing the electrolysis reaction in the Hub. Such a fuel management is compatible with the mission operations and less than a few hundred watts would be required for the thermal cooling of the oxygen and hydrogen (already included in the 40% efficiency of the system).



page 258 of 317

s

Figure 7-26: Principle of the regenerative fuel cells (RD[68])

Compared to secondary batteries, regenerative fuel cells become very attractive for large-scale energy storage purpose. But, since the power generated by the fuel cells is almost constant, an additional module is required for:

• Supplying the peak power requirement on the bus • Absorbing the excess of power produced by the fuel cells

A Li-ion battery able to provide 6 kW for 3 hours fit with these requests. Such a battery can also provide all the power required when the Hub is in the Moon’s shadow: the fuel cell system would then only be used for safety purposes and for the Earth occultation phases. As for the LEV fuel cell design, the model issued from RD[66] has been used here. Note that the masses of the auxiliary units (gas driers, liquefaction units, radiators…) are included in the mass of the tanks. Table 7-71 shows the first assessment (for 6-hour Earth occultation and 6-hour safe mode) of a design based on regenerative fuel cells (and with rigid solar panels for easy comparison with the battery option). 30% has been added for the failure of one of the fuel storage tanks.



page 259 of 317

s

Table 7-71: Power subsystem sizing for regenerative fuel cell architecture

7.10.3.2.3 Flywheels

Flywheels are mechanical batteries that convert energy into a mechanical motion and when required, convert that motion back to energy. Their roundtrip efficiency is extremely high (85 to 95%) with a lifetime estimated longer than 20 years and they also have an attractive specific energy, even higher than secondary batteries. A demonstrator turning at 60 000 rpm and is able to store up to 7.5 MJ (see RD[67]). The main disadvantage for this mission is the important self-discharge: the NASA demonstrator is completely discharged after only 12 hours. Flywheels are rejected for this mission including shadowing phases of 6 hours (that could be immediately followed by a 6-hour attitude pointing failure).

7.10.3.3 Power conditioning and distributing

A closer look at the architecture of the conditioning and distribution is not adapted to the limited definition resulting from this feasibility study. Nevertheless a topology with a regulated bus of 120 V is selected to limit the contributions of the dedicated DC/DC converters and to limit harness losses. An interesting topology that optimises the power losses of the mission would be based on the use of S3R or S4R modules (see Figure 7-27):



page 260 of 317

s

Figure 7-27: S4R and S3R regulated bus architecture

For example, to handle the reliability requirement, three separated buses could be connected to the critical units. The conversion efficiency of the power conditioning and distribution functions are assumed to be as followed:

• From the energy storage to the users: 90% (BDR) • From the bus to the energy storage module: 90% (BCR) • From the PVA to the bus: 90% (Shunt module)

7.10.3.4 Trades conclusion

The best candidates for the power storage and the power generation were selected: • For the power storage: use of regenerative fuel cells or Li-Ion secondary batteries • For the power generation: either thin-film cells or rigid GaAs improved cells

As regards power storage, Table 7-70 and Table 7-71 show a mass impact for the battery architecture system compared with the regenerative fuel cell option of around 900 kg on the power subsystem level. Even decreasing the solar array area required in the battery storage option (122 m2 instead of 167 m2) cannot counterbalance the increase in mass. The regenerative fuel cells option therefore seems the most attractive option for the Hub requirements. However, since the battery package should fit in around 1 cubic metre and since the mass difference is less than 2% of the Hub total mass, the battery option is selected for the baseline design. For the time being, this conservative approach is preferred and is more in line with the current development status of the storage technologies. Note that important R&D investment for fuel cells needs to be started soon. In addition, the 10 year lifetime mission requirement will have also to be proven. In the context of this study, using regenerative fuel cells would also facilitate the implementation of fuel cells inside the LEV: the Hub could become a hydrogen/oxygen liquefaction plant for recharging the LEV power subsystem for a new mission. As regards power generation, both options have been kept for the baseline design assessment.


7.10.4.1 Preliminary mass budget

Table 7-72 shows a sizing budget for both photovoltaic generation options. The different power subsystem are sized taking into account the following points:



page 261 of 317

s• Battery:

o Energy requirement: a 6-hour Earth occultation followed by a 6-hour attitude control failure mode

• Power conversion and distribution: o Without unduly long discussion on the design of the electronics, an arbitrary average

value of 11.1 kg/kW is assumed for this preliminary computation • Solar arrays:

o For both options, the requirements are identical o The PVAs are sized for providing the energy required on the bus during the daylight

and for completely recharge the battery discharged after a 45.2 minute lunar eclipse o Two wings with the system sized for one (redundancy) o Based on RD[52] and RD[53], the recurrent costs have also been assessed

Table 7-72: Baseline designs computation

7.10.5 List of equipments

7.10.5.1 Battery module

The battery module is based on the average power requirements of the Habitation module during 12 hours at the end of life. A more accurate assessment of the power consumption during the safe mode should lead to a smaller battery module. On top of the computed mass, an additional maturity margin of 20% is considered since enhancements of the technology were taken into account.

7.10.5.2 Solar arrays

In the preliminary budget assessment (Table 7-72), the thin-film cell options give a limited reduction of the PVAs total mass. The only remaining advantage of the thin-film option is the recurrent cost that should be reduced by a factor of six or seven. Taking also into account the uncertainties of the non-recurrent cost of the thin-film cells, the technology readiness and the timeline of the mission, the size of the solar arrays to accommodate, the rigid GaAs option is



page 262 of 317

skept as baseline for the Hub design. Thin films can be envisaged as an option depending on the improvements reached in the upcoming years.

7.10.5.3 Summary

Including the maturity margin of the equipments, the power subsystem should weigh about 2 tonnes and excluding the solar panels it should take less than 2 m³.

Table 7-73: List of equipments

7.10.6 Conclusion

An optimised preliminary configuration for the power subsystem has been selected. The reliability is high since qualified technologies have been preferred with expected performances enhancements for 2015. The development and qualification of the regenerative fuel cells technology would have a positive impact on the Hub spacecraft’s power subsystem. Further additional mission requirements or better knowledge of the different subsystems may increase the power requirements of the Hub. The higher the power needs become, the higher the benefit of the regenerative fuel cells option.

7.11 Hub – thermal 7.11.1 Requirements and design drivers

The following set of main system requirements has been used in the thermal design of the Hub: • Moon orbit: polar 100 km circular orbit • Vehicle lifetime around the Moon: 10 years • Crew of six people maximum • Attitude: three-axis stabilised Nadir-pointing • Two-module design with an inflatable habitation module • Temperature within habitable zones between 18 and 27°C, relative humidity 25 to 70% • Reference configuration: up to four vehicles docked to the Hub, corresponding to the

worst case of radiator shielding and maximum power dissipation for the Hub. One important design driver is obviously safety. However, concerning safety, this specific application is not expected to differ substantially from the design of Habitation Modules in Earth orbit. Therefore, all the design principles used for ISS modules are directly applicable. A second peculiar design driver is the Moon orbit environment and, in particular, the high infra red fluxes from the lunar surface at and around subsolar point. This imposes some independent control of attitude for the radiators to avoid parasitic heat.



page 263 of 317

s7.11.2 Assumptions and trade-offs

To size the subsystem a worst-case analysis has been performed. Several cases have been analysed varying the orbital node. The worst hot case resulted in the dawn/dusk orbit (Sun perpendicular to the orbital plane and no eclipse), which corresponds to maximum solar heat absorbed by the external surface of the vehicle at end of life. In this case the IR flux absorbed from the planet is minimum but the total net heat flux into the spacecraft is the highest. Figure 7-28 shows an example of the thermal cases analysed.

Figure 7-28: Thermal case examples


The baseline thermal control design is based on single-phase active cooling for the cabin and equipment of each module. Each active cooling system is independent and it is made up an external loop linked to the radiator and an internal loop linked to the cabin and equipment cold plates. The thermal control systems of the two modules are linked together via a “thermal bus” giving some redundancy (Figure 7-30).



page 264 of 317

s

Figure 7-29: Schematic of the Hub Active Cooling System

Figure 7-30: Link between module thermal control systems

Radiator assembly

N

NN

N

N

XCAassembly

Liquid/liquidheat exchanger

Propulsioncompartment coil

Flow control valve

Pumpassembly

Pumpassembly

Cold platesassembly

Pumpassembly

Internalshell coil

N

Thermal bus

N



page 265 of 317

s

Large (total of 155 m2 surface area) deployable radiators are required. Due to the conceptual nature of this study, a simple two-wing configuration has been selected.

7.12 List of equipment

Figure 7-31: Hub thermal equipment list


1 HUB module / radiator 2 770.00 10 1694.02 HUB module / coil assembly (HP 4 16.90 5 71.03 HUB / insulation 1 100.00 10 110.04 HUB / HCU 3 6.00 10 19.85 HUB / Heaters, thermostat, lines 2 13.00 10 28.66 HUB / oxygen coolers 4 3.00 20 14.47 int. loop / liquid 2 104.00 5 218.48 int. loop / dry tubing + insulation 2 120.00 5 252.09 int. loop / pump assembly 4 14.00 10 61.6

10 int. loop / compensator 2 3.00 10 6.611 int. loop / cooler-dryer assembly 4 20.60 10 90.612 int. loop / cold plates 10 3.40 5 35.713 int. loop / hydraulic connector 20 0.30 5 6.314 int. loop / choke washer 20 0.50 5 10.515 int. loop / on-off valve 20 4.00 5 84.016 int. loop / by-pass valve 2 5.00 10 11.017 int. loop / manual valve 20 4.00 5 84.018 ext. loop / liquid 2 36.00 5 75.619 ext. loop / dry tubing + insulation 2 4.20 5 8.820 ext. loop / pump assembly 4 14.00 10 61.621 ext. loop / cold plates 10 3.40 10 37.422 ext. loop / compensator 2 3.00 10 6.623 ext. loop / heat exchangers 8 15.90 10 139.924 ext. loop / flow control valve 2 5.00 5 10.525 ext. loop / on-off valve 10 4.00 5 42.026 ext. loop / choke washer 10 0.50 5 5.3- 20 0.0


MASS [kg]Element 1: Habitation ModuleMargin Total Mass

incl. marginMass per quantity

excl. margin


Unit QuantityElement 1 Unit Name



page 266 of 317

s



page 267 of 317

s8 GROUND SEGMENT AND FLIGHT OPERATIONS

8.1 Requirements and design drivers The Ground Segment and Flight Operations infrastructure shall be able to support the following operational tasks:

• In-orbit assembly (LEO and LLO) • Lunar orbit transfer and insertion • Rendezvous and Docking manoeuvres (LLO) • Lunar Descent, Landing and Ascent • Earth transfer insertion • Earth re-entry • Operation of subsystems (including Life Support) • Support of crew activities (mission and personal)

The dominant implications of a manned mission are that of human safety (psychological as well as physical) and of the size of the communication links required by a modern day mission for state-of-the-art video, voice, e-mail, HKTM, science data and OBSW images. A manned mision has several critical activities (beyond launch) of which the rendezvous and docking manoeuvres (RDM), in-orbit assembly and safe mode recovery of the unmanned elements are only mission critical, while the others have the added burden of being human-life critical:

• RDM of the CTV to LTO PMs in LEO • RDM of the CTV to the Hub in LLO • RDM of a cargo vessel or new LEV to the Hub while it is manned • Lunar Descent and Landing • RDM of the Ascent Vehicle to the Hub in LLO • Earth re-entry of the CTV • Safe-mode recovery of a manned vehicle (although, in principle, there will be no

“dramatic” safe-mode of a manned vehicle – all non-essential systems cannot necessarily be switched off)

This last point implies a high level of complexity, intelligence and automation of the on-board systems of a manned vehicle. The Operations assessment considers the architecture of the flying mission as being made up of four regions the operations for which will need to be co-ordinated:

• Launch – up to separation • LEO Operations for Assembly • Earth-Lunar-Earth Transfers & LLO Operations • Lunar Surface Operations

Operations in the different regions are not mutually exclusive and will happen in parallel. Worst case round-trip communication times to the Moon are less than 3 s, so NRT operations from the ground are possible.



page 268 of 317

s The following requirements have been defined for Ground Operations:

• There shall be different Control Centres for each of the “operational regions” defined above.

• There shall be support centres for: the training and care of astronauts throughout the entire lifetime of the mission including in flight; engineering support where full-scale engineering models are maintained; public relations; overall mission management; and science mission co-ordination. These centres need not be at a unique location.

• There shall be a 24-hour core human presence on the ground, though the working day shall be set to match the rhythm of the crew, with reaction times of less than 2 hours to mobilise the full ground support team and International Partners in the case of a critical contigency.

• There shall be a worldwide ground communications network capable of providing full-time TT&C, science data reception and distribution, and voice and video communications at a rate suitable for a manned mission. A maximum downlink rate of 100 Mbps has been defined.

• The ground communications network shall be fully redundant and capable of simultaneously supporting the maximum data rates of each of the above “regions” during the full extent of parallel operations.

• Critical operations shall have a guaranteed 100% communications visibility and reliability from the Earth. For all the communications links used throughout the different phases of the mission there will be requirements on the availability, integrity and reliability of the link.

• The Ground Stations sites selected for the Space communication links shall have X and Ka+ band capabilities. There shall be (a minimum of) four sites around the world to provide full-time communications visibility of the Moon from the Earth (Ka+ band communications are very vulnerable at low elevations so three equidistant sites are unacceptable).

• The Ka+ band link selected for the high rate data is highly susceptible to weather conditions so redundancy on a lower performance level is required. Communications shall be organised hierarchically. All real-time monitoring and control TT&C channels, a voice link to ground and an e-mail channel shall be the minimum available on a back-up X-band link or with an increased signal-to-noise ratio on the Ka+ band link in case of bad weather.

• Ground control shall have the following orbit tracking capabilities: Doppler; Ranging; GPS or Galileo in Earth orbit; Delta DOR; Same Beam Interferometry; Satellite to satellite Doppler and Ranging. The accuracy of these methods shall be at least the state of the art projected for 2010 (deltaDOR: 5 nrad, Doppler and Ranging as BepiColombo Radioscience (0.6*10-15 Allan variance and 10 cm ranging ground station contribution)).

• Early warnings shall be provided for solar flares to give the crew sufficient time to reach the safety of a shielded compartment. The lunar surface crew have to return to the Hub. This might imply that improved space weather monitoring capabilities will have to be improved.

The following requirements have been defined for on-board operations:



page 269 of 317

s• For a manned vehicle there shall be a graceful degradation of on-board systems at times

of anomaly. There shall be no single event or combination of two foreseeable events that could lead to a reduction in the quality of life on board beyond a TBD margin. The level of robustness required for a manned mission beyond LEO will be much higher than for any of today’s missions.

• “The vehicle shall provide the flight crew on board the vehicle with proper insight, intervention capability, control over vehicle automation, authority to enable irreversible actions, and critical autonomy from the ground.” (as cited from RD[8])

• “The flight crew shall be capable of taking manual control of the vehicle during all phases of flight” (as cited from RD[8])

• The crew shall have the resources to perform hardware and software repairs. The list of tools shall include a stand-alone simulator of the vehicle with emulated flight software.

• A manned vehicle shall be able to survive without ground contact for as long as the maximum lifetime of the on-board perishable resources essential for the support of human life.

• The crew shall have on-board facilities to maintain attitude and orbit control and to plan and execute a return to the Earth without assistance from ground control.

• A vehicle without a crew on board shall be able to maintain its attitude and orbit and execute a timeline or survive following FDIR intervention for a period of 21 days without ground contact.

• Ground control shall be able to take full control of all vehicle functions and subsystems down to switch level.

• The crew shall have ample access to private communication channels allowing e-mail interactions, internet for personal, health, religious and work interests, and a voice and video link to family and friends. This requires a secure firewall and a minimum 10 Mbps link (the equivalent of a modern-day LAN connection in a business organisation).

8.2 Assumptions and trade-offs When defining the Ground Segment architecture, the following assumptions have been made:

• There is a main Control Centre (CC) for each of the operational regions defined above. • Launch operations will be performed by currently existing infrastructures such as

Arianespace. Thus, the Launch region is not considered further. • The LEO Control Centre (LEO-CC) will have a minimum life of 15 years including

operations definition and preparation phases but excluding design and construction of the site. The centre has the operational control of a vehicle following launch separation up to final commissioning and clearance for TLI.

• The LEO-CC will control both assembly lines in LEO. • The Earth-Lunar Control Centre (EL-CC) will have a minimum life of 14 years including

operations definition and preparation phases but excluding design and construction of the site. The centre has the operational control of a vehicle once it is cleared for TLI.

• The Lunar Surface Control Centre (LS-CC) will have a minimum life of 12 years including operations definition and preparation phases but excluding design and construction of the site. The centre is operationally responsible for the lunar surface mission between the crew departure from the Hub and their return.



page 270 of 317

s• Apart from at the LS-CC which does not have year-round real operations there are no

large fluctuations in workload for the personnel at the CCs so there is no problem foreseen in managing staffing levels.

• It is assumed that the LS-CC and the Astronaut Support Centre will be the same facility (if not necessarily at the same location) with a sharing of personnel and infrastructure between training & preparation phases and actual surface operation phases.

• The LS-CC will also be responsible for crew launch up to clearance for TLI and Earth re-entry operations with post-flight co-ordination, thereby removing any requirements for the LEO assembly teams to have to be qualified in manned spaceflight operations.

• There will be multiple Engineering Support Centres (ESCs) but a single entity for public relations.

• 12-m, Ka+ band antennas have been defined for the ESTRACK sites at Vilspa (Madrid), Kourou and Perth. These sites currently have 15-m, X-band antennas and it is assumed that either new antennas will be built or the existing facilities will be upgraded as part of the regular ESTRACK upgrade programme.

• Either an ESTRACK site will be established to fill the “Pacific Gap” or one of the antennas at the Canberra Deep Space Communications Complex will be upgraded.

• It is assumed that all stations of the ESTRACK network will be upgraded to at least X-band transmit and receive and made available as back-up.

• It is assumed that the MSM Communications Infrastructure1 set up to support the current European Manned Spaceflight Operations will be available (or its equivalent) for the lunar exploration mission.

• It is assumed that the ESTRACK network and co-operating sites will be given permanent connections to the Interconnection Ground Subnet.

• It is assumed that the TDRS system will be used for communications in LEO. • Two communications relay satellites will be placed in a halo orbit around the Earth-

Moon L2 point to provide full time communications with the LLO Hub. The operations concept surrounding the LEO assembly of the various vehicles makes the following assumptions:

• There will be multiple modules in orbit at any one time which all must be monitored, tracked and controlled.

• Ground segment automation will be widely used to reduce the operations load on the ground. This involves management of the communications links, monitoring and limit checking of the telemetry, alarm raising and notification of the necessary support personnel at times of anomaly, orbit and attitude determination, monitoring and control and mission planning (bearing in mind that the assembly mission is repetitive and will have a mostly fixed sequence).

• Spacecraft state of health beacons will be used to reduce the load on Ground Station and TDRSS requirements. A low energy/low data rate beacon is continuously transmitted to

1 The MSM-CI consists of the Interconnection Ground Subnet (IGS) and the MSM data, voice and video services. The IGS is a combination of ATM VPN and ISDN networks which provide permanent, scheduled and on-demand transport services. RD[9] Note the status of the IGS Phase I which will terminate in May 2005. Phase II will incorporate SLE and TCP/IP technology.



page 271 of 317

sthe Earth, the modulated frequency of which reflects the health of the spacecraft. A system of small and low-cost antennas to autonomously track, receive and forward the signal will have to be established around the world. This could be done gradually as part of the ESTRACK upgrade programme and also used by other future missions.

• Beacons require on-board intelligent systems that can autonomously analyse the spacecraft state of health. It is assumed that the development costs of such a design can be shared with the development costs of the LEO-CC Mission Control System that will perform a complementary monitoring and analysis function.

• The goal of having no direct human involvement in the assembly process has been assumed. Any sort of EVA has an element of risk but this is increased on a construction line; and as the assembly will be on going for the entire lifetime of the mission, a two-fold, (semi-)permanent human presence in LEO would be a considerable extra cost burden for the mission.

• Assembly will be by autonomous docking only. Wireless on-board data handling will allow the virtual integration of the new module to the growing vehicle. This technology is still under development for space applications but Ultra Wide Band systems on ground today allow the wireless transfer of data over a 10-m distance at a rate of several hundred Mbps.

• The alternative to assembly by docking or assembly “by hand” (from a constructions platform) is to use robotic telepresence (from a constructions platform) with the operator located on the ground. It is assumed that the development costs of the docking interfaces would still be less than the costs of developing, manufacturing, operating and maintaining two construction platforms in LEO.

• The Propulsion Modules (PMs) have to be their own spacecraft with AOCS, Power, Thermal, Communications and Data Handling subsystems. This allows for the adoption of the Operations tools already developed for (pseudo-)autonomous dockings: Ground Operator Assistant System (GOAS) and Remote ATV Control System at ISS (RACSI), RD[10]. Both of these tools allow the operator on ground (GOAS) or in an orbiting element (RACSI) to order a Collision Avoidance Manoeuvre.

• There will be a LEOP and (short) Commissioning Phase for each module launched. • The CTV PMs need to wait for the CTV to be launched, however, they can dock with

each other in the meantime. The other PMs can dock immediately with their ‘parent’ module (approximately 48 hrs after launch) which will already be in orbit waiting for them. This keeps the number of objects on the assembly line that have to be tracked, monitored and controlled to a minimum. The orbital positioning by on-board GPS or Galileo receivers will be transmitted to the ground as part of the routine TM.

• Once modules are docked, they will operate as the subsystems of one large module, i.e. they will no longer act (solely) on their own FDIR, the master module will have an FDIR for the entire assembly and the LEO-CC controls just one spacecraft.

As regards robustness of design, the following has also been assumed:

• The PMs for the CTV will be man-rated, i.e. have the Fault Tolerant (FT) design commensurate with the criticality of human life.

• Unmanned vehicles that are to be used within the sphere of human involvement will be designed and built with the same FT philosophy in mind as for manned vehicles, i.e. PMs for the LLO Hub, and PMs for the cargo ship and LEV that are jettisoned during LTO



page 272 of 317

scan have a simpler, more conventional safe mode mechanism, while those used for LOI will need to be of the same design as the CTV PMs. The master FDIR would adapt itself according to the design of the elements under its surveillance.

The point of this distinction is cost of manufacture and the requirements put on the CCs to be prepared and trained in operations for conventional spacecraft and/or for man-rated designs: the LEO-CC will only need to be qualified for conventional operations, the LS-CC only for manned flight operations and the EL-CC for both types.

• The principles of a robust design will be followed including: mutual monitoring of distributed processes; majority voting; multiple levels of sub-unit redundancy offering hot and cold back-up of the most critical units with self-analysis and switching; and layered OBSW.

• Advanced technologies will be used: intelligent sensors with fuzzy logic connections; artificial sensors and intelligent consistency checks based on intrinsic redundancies enabled by data fusion; intelligent FDIR that uses trend analysis to forewarn and reconfigure; and adaptive control that optimises on-board resources.

8.2.1 Autonomy

RD[8] introduced the requirement of crew “critical autonomy from ground” and this assessment has proposed the minimum requirements that would allow a crew to safely return to the Earth without ground support. In connection with this is the consideration of (crewless) vehicle autonomy. A survival period of 21 days without ground intervention has been proposed after a consideration of current day capabilities (e.g. SMART-1 has a MTL stretching to 10 days in the future). In addition, the ever increasing capacity of OBSW and processing power is seeing ever more “intelligence” being developed for new missions, although there is always a trade-off of automation development costs against the perceived risk and predicted reduction of operations costs. It is assumed that the redundancy of the TT&C subsystem on board and the communications network on the ground provides a full-time link that would, in the case of a unit failure, reconfigure and re-establish itself in a matter of minutes. This type of link can be used to allow ground and flight systems to continuously ‘talk’ with each other to handle routine operations. Where present, humans can be used as advanced FDIR processes but never as the first line of defence. They cannot be expected to be in a permanent state of alert nor to have the detailed information base and split-second reactions, and they should not be used as an excuse to not apply modern-day software engineering practices to space segments. It is assumed that while the crew need to have their critical autonomy from ground from the very first manned flight, ever more advanced vehicle autonomy will be phased in over the first half of the flight-programme, progressively reducing operational load for the humans in orbit and on the ground. Consequently, the costs of autonomy development can, in part, be offset by the reduction in operations support that can be made during the life of the mission as the autonomy is brought in. Besides, such a mission would be defining the baseline for inter-planetary missions to follow and the on-board autonomy would be one of the most significant and re-applicable developments to



page 273 of 317

sbe passed on. The cost benefits should therefore be planned out over decades rather than just over the life of the project. Because each CC (LEO, EL and LS) will spend time controlling each spacecraft, it is assumed that the different spacecraft will be designed to have the same ‘look and feel’ when it comes to operating them from ground. Thus, it is assumed that the core of the Mission Control System (MCS) at each location will be delivered from a single source and be configurable by the different operations preparation teams. This should mean a big cost saving to the programme. It is also important when sizing the Flight Control Teams. If there are multiple spacecraft, which are essentially the same to fly, a separate FCT for each vehicle is not required. It is assumed that the development of the MCS and the flight software will be closely inter-related or even under the same initiative. This will be to reduce the perceived risk element of on-board autonomy development by allowing the prototype software to be tested as part of the MCS against the NRT operations running in parallel. After a period of qualification, the software can be introduced to the next generation of vehicles and the MCS progressively updated. However, it is assumed that developments in on-board autonomy will have little impact on the assembly operations the autonomy of which will already be at a suitable level by the beginning of the flight mission.

8.3 Baseline design The Ground Segment will consist of the following elements (see Figure 8-1):

• Launch Control Centres (LA-CC) – there will be at least two facilities to service the two LEO assembly lines.

• LEO Control Centre (LEO-CC) – responsible for both assembly lines. • Earth-Lunar-Earth Control Centre (EL-CC) – responsible for LTO, LLO and ETO

operations. • Lunar Surface Control Centre (LS-CC) – responsible for the lunar surface mission, crew

launch and crew re-entry, works with the ATC in the training and preparation of the astronauts.

• Astronaut Training Centre (ATC) – responsible for the monitoring of the crew health in-flight, works with the LS-CC in the training and preparation of the astronauts.

• Mission Management Team (MMT) – overall coordination and responsibility for the lunar exploration mission, responsible for the allocation and scheduling of the IGS resources.

• Science Coordination Team (SCT) – acts as a bridge between the mission and the science community represented by the USOCs.

• User Support and Operations Centres (USOC) – science, industry and academic centres that provide experiment and instrument modules.

• Engineering Support Centres (ESC) – maintain full-scale engineering models that reflect exactly the current status of the in-flight equipment, i.e. every TC that is sent from ground or issued on board is also received by the appropriate ESC.

• Public Relations (PR) – acts as the bridge between the mission and the public. • Ground Stations – 12-m X and Ka+ band capable antennas at Perth, Vilspa, Kourou and

Canberra. Other stations of the ESTRACK network are available as back-up. The



page 274 of 317

sschedule for ground station usage is provided by the ESTRACK Management System (EMS) and passed to the MMT for incorporation into the IGS schedule. The stations are time-shared between other missions and the EMS is the central coordination service that guarantees that the communication requirements of each one are met.

Figure 8-1: Lunar exploration ground segment

Three main control centres are considered in further detail.

8.3.1 LEO Control Centre

This is concerned solely with the assembly of all vehicles in low Earth orbit, which is a series of small, repetitive missions with limited variation throughout the 10 years of flight time. The centre takes over after separation from the launcher and performs an intensive period of LEOP and Commissioning over the next 48 hrs until the module is docked with other waiting modules or parked in a holding orbit if it is the first module of its vehicle. Shift work is required at this time. The TDRSS infrastructure is used for the period up to docking/parking after which a state of health beacon is used to keep a routine check on the growing spacecraft with occasional ground station passes for TT&C. After a vehicle has been completed, it is given a final commissioning to confirm that the assembled structure is acting like a single vehicle with all subsystems performing within limits. If this is the case, it is cleared for TLI and handed over to the EL-CC. The only restriction placed by this centre on the mission is that modules for the two different assembly lines are not scheduled to be launched within [launch window + 2 days] of each other (two days for launch until docking/parking). There is one Flight Control Team (FCT) for the DM assembly and one for the AM assembly, the members of which act as back-up to their counterpart in the other team. The AM team is also responsible for the assembly of the CTV



page 275 of 317

sPMs. The assembly of the LLO Hub to its PMs is a one-time activity and can be handled by the DM and AM teams. The mission planning engineers are responsible for launch and resource coordination. For the first round of assembly operations there are industry support teams on-site. After this, they are on-call from their own sites.

8.3.2 Earth-Lunar Control Centre

This centre has a large remit and is logically ordered into LTO operations, LLO operations and ETO operations. There are two fully separate Flight Control Teams: one for the Transfer Orbits and the other for the Low Lunar Orbit. The members of the teams act as back-up to their counterpart in the other team. Shift work is required at times of orbit insertion, rendezvous and (un)docking. In addition to these two teams is the smaller Crew Operations Team (COT) which works directly with the crew in-orbit. Only the COT is required to change its working day to match that of the crew with 4 days on, 4 days off shift work between the two subteams.

8.3.3 Lunar Surface Control Centre

Apart from the lunar surface mission, this centre is also responsible for crew launch (until docked with the CTV PMs) and re-entry operations, and crew training in co-operation with the ATC. There is a single FCT and a single COT each with internal redundancy. The COT is required to match its working day to that of the crew for the duration of the lunar surface mission with an internally cycling shift pattern that uses the redundancy in the team. The FCT has to support the launch, re-entry, lunar descent and ascent with shift work. At this site is the solar flare warning system for the entire mission. A Space Environment Information System (SEIS) has already been developed at ESOC that provides forecasts based on historical and current data. Common to each of the control centres are the following:

• Mission Control System – although not identical in each centre, the core of each MCS is an operating system that allows the user to switch between the monitoring and control of different modules/spacecraft while always performing limit checks on all the received TM.

• Software simulators of the individual modules that can be assembled like the real things. Used for training, testing and validation.

• Flight Dynamics Support (FDS) for manoeuvre planning, rendezvous and docking, and routine control. A study should be performed to determine if FDS at each site is really required or whether a common FDS infrastructure redundant over two sites is actually sufficient.



page 276 of 317

s



page 277 of 317

s9 RISK

The scope of the risk analysis is to identify risk scenarios that can potentially jeopardise the lunar exploration mission, to identify the risk causes and consequently the risk control to eliminate, minimise or control the risks. The risk analysis should be further developed during the project definition to analyse all the system, refine the risk identification and classification, and provide evidence that all the risks have been effectively controlled. The risk assessment process follows the steps defined in the ECSS-M-00-03A Risk management as below:

1. Identification of hazardous/failure conditions. 2. Identification of failure scenarios and their consequences. 3. Categorisation of the scenarios according to their consequence. 4. Analysis of likelihood and uncertainties of risks. 5. Identification and ranking of risk contribution of individual scenarios.

The consequence severity level is defined according to the worst-case potential effect. As regards mass, cost and schedule, the severity should be defined according to the risk of overcoming an established target or a requirement. The consequence severity level is defined according to the worst-case potential effect as shown in Table 9-1. Each risk scenario has been evaluated in terms of likelihood and consequence severity on performances, identifying the impact on the mission characteristics. The judgement shall be iteratively based on the analysis of each risk scenario cause, developed along the project life. Severity Definition Catastrophic Loss of life

Loss of system/mission with impact on safety Long-term detrimental environmental effects

Critical Loss of, or major damage to flight system elements, or ground facilities. Loss of, or major damage to public or private properties. Short-term detrimental environmental effects.

Major Degradation of system/mission. (Failure propagation, but system able to control the consequences).

Minor Degradation of system/mission with no failure propagation. Negligible No/minimal consequences.

Table 9-1: Consequence severity

9.1 Mission success definition, requirements and assumptions Human spaceflight statistics show a 5% of risk of losing the crew. Any next-generation system for transporting people to the Moon will be probably designed to a risk requirement much lower than that e.g 1%, according to RD[8]. The programme shall be designed so that the cumulative



page 278 of 317

sprobability of safe crew return over the life of the programme exceeds 0.99. Besides, a crew escape system shall be provided on Earth-to-orbit vehicles for safe crew extraction and recovery from in-flight failures across the flight envelope. The escape system shall also have a probability of successful crew return of 0.99.

9.1.1 Mission success criteria

The following mission success criteria have been defined. The mission is considered to be successful when all scientific objectives are realised safely. The mission is considered to be lost when either the crew is lost or the main objectives of the mission cannot be performed. The applicable requirements are:

• The preservation of human life (general public, ground personnel, flight crew) shall be the most important priority in the development and operation of space systems’ (ECSS-Q-40 Safety):

• Double and single Failure/Fault/Operator error tolerance to catastrophic & critical events. • Safety goal: to identify all possible safety hazards, to eliminate/control them to an

acceptable level during all the phases of the mission The assumptions are:

• Earth Operations and SW risks are not assessed. • Legal & Programmatic risks are not assessed. • Heritage data will be taken from Apollo missions.

9.2 Preliminary results The main uncertainties are linked to the need for new technologies to perform the mission successfully. As a result of the risk analysis, the following measures are incorporated in the design. The risk categorisation is shown in Table 9-2.

• Abort modes are provided in most cases to safely recover the crew and vehicle/s or permit the use of an escape system. For more information, see Chapter 4, Mission Architectures.

• All critical systems essential for crew safety are designed to be two-fault tolerant. Hazards not controlled by conformance to failure tolerance are controlled by ‘design to minimum risk’ (i.e.: fracture control, safety factors, containment, material selection, sharp edges avoidance, and operations assessment.)

• Maintainability is considered as an option for the ECLSS. • Provision for caution and warning systems. They are defined so as to remain operational

during power failures or other anomalies. • Fire detection, control and cleanup system included. • Depressurisation detection system included. • Contamination detection and cleanup system included. • Contingency consumables foreseen. • Humans-in-the-loop:



page 279 of 317

so Crew will have control capability over vehicle automation, critical autonomy from

ground. o Crewless vehicles will permit safety-critical commanding from the manned vehicle. o Protection against space radiation hazards via safe havens. o Considerations have been made of effects of changes in gravity forces and

physiological / psychological risks from extended confinement and hazardous operations.

o EVA safety is taken into consideration. Suits present characteristics are not safe for radiation or meteoroids hazards.

Risk acceptability Risk domain and scenario Status Unacceptable • Maximum likelihood with catastrophic

consequences: o Radiation hazards o EVA suits inadequate to

environment o Operations hazards o Cryogenics failure o LSS failure o Ariane-5 failure (man rated)

• Numerous critical areas with uncertain environment definition

• Research level only • New project beyond the

state of the art • New processes / Need for

new facilities • High level of autonomy

required for operations Acceptable if reduction impossible

• High likelihood with critical consequences: o Communications loss

• Qualified technologies but never applied in projects

• Numerous modifications of qualified product

Acceptable

Other Defined environmental conditions, qualified products, existing processes and facilities.

Table 9-2: Preliminary risk categorisation

Note that the arrow shows the normal evolution of the risk during the development of the programme. Orange is for high contribution, grey for medium contribution and white for low contribution.

9.3 Recommendations It is recommended to investigate how to minimise the critical items by considering:

• Extensive mission abort/ rescue capabilities. • Greater reliability and/ or redundancy of systems. • Preventive and/or corrective maintenance strategy (consider the use of robotics) • Capability to monitor/ detect and assess effects of slow events such as:

o Metal fatigue, cracks; dust, corrosion and rust o Cabin atmosphere toxicity, contaminant and hazardous substance concentrations are

potential toxic threats in the recycling of breathable habitat atmospheres, water recycling systems, and solid waste handling and recycling systems; bio-hazards, deterioration of electrical insulation of wires; thermal insulation; seal deterioration, food spoilage, potable water contamination…etc



page 280 of 317

s• Protection against space radiation hazards; several effects of changes in gravity forces and

sociological/psychological risks of extended confinement and hazardous operations. EVA safety. Safe Haven.

Risks can be reduced but not eliminated. There always remains the chance of an “unknown” risk. Measures that reduce risk are often under pressure for elimination because they cost money. These results are of indicative nature but can be used as a first risk reduction tool. The risk will be reduced when the necessary new technologies will be developed and qualified. The risk analysis should be further developed during the project definition to analyse all the systems, refine the risk identification and classification, and provide evidence that all the risks have been effectively controlled and reduced to an acceptable level.



page 281 of 317

s

10 PROGRAMMATICS/AIV

The large effort necessary to develop and operate a system capable to periodically land men on the Moon and safely return them to Earth requires a wide industrial consortium. As regards the development, manufacturing and delivery, it will be challenging to procure several manned pressurised elements in a reasonably short time. The capability of industry to manufacture, integrate and test in parallel those elements is a key to the success of the project.

10.1 Development 10.1.1 Lunar Excursion Vehicle (LEV)

The LEV is composed of three modules: the Descent Module (DM), the Surface Element (SE) and the Lunar Ascent Vehicle (AV). The Ascent Vehicle and the Surface Element will be manned pressurised modules. The DM is needed to safely land the LEV on the lunar surface, but will not have habitable volumes for the astronauts. As a complete and autonomous spacecraft, the LEV shall have automatic and manned flight control capability. This implies that the qualification flights to plan for this vehicle will include manned missions, with positive demonstration of the capability of the vehicle to be piloted by the on-board astronauts.

10.1.2 Hub

The Hub is composed of two pressurised modules and provides five docking ports. It will be equipped with a propulsion subsystem for attitude control and docking manoeuvres, allowing in-flight assembly operations. A module will be inflatable. This technology is not available yet, but it is supposed it will be at the needed time.

10.1.3 Propulsion Module (PM)

A number of propulsion modules shall be procured to enable the delivery of the Hub and the LEV to LLO, from LEO.

10.1.4 Cargo vehicle

A cargo vehicle is needed to support the logistics of the Lunar Exploration Project. The cargo vehicle should periodically resupply the Hub. An estimate was accomplished within this study showing the possibility to develop such a vehicle from a modified and improved ATV. If it is found feasible in the next project phase with a more thorough study, the development cost of the cargo vehicle would be reduced.

10.2 Ground facilities and centres A number of test facilities, training tools and ground support centres have to be available to support the full process of development, verification and mission exploitation. The associated costs are very high. It is unthinkable to build everything brand new, nor is this always necessary. In the next phase, when the system design will be more detailed, it shall be evaluated which



page 282 of 317

sfacilities are already available in Europe or in other countries, to provide the requested support to the verification and operations of the system, and which facilities need to be built.

10.3 Environmental testing Specific environmental tests are required to qualify the many elements of the system, and then to demonstrate their flight worthiness on the ground.

10.3.1 Hub

If the ISS approach is taken, no thermal vacuum/thermal balance test is required of the Hub pressurised modules to demonstrate qualification of the vehicle and later on its acceptability for flight. Nevertheless, acoustic testing is required for each module. The sine vibration test is optionally done on a dedicated structural model, depending on the characteristics of the launch vehicle and on its requirements. Due to the need to inject the Hub into a transfer orbit to the LLO, by using propulsion modules, a random vibration test both in qualification and in acceptance may be required. The necessity for this test shall be evaluated on the basis of the as-built PM characteristics. A full qualification model (QM) of the Hub active thermal control (ATC) and of the Environmental Control and Life Support (ECLS) subsystems is deemed necessary, to demonstrate the functions and performances of these subsystems in their integrated configuration on the Hub.

10.3.2 LEV

Thermal vacuum and thermal balance tests are required to qualify and accept the modules of this vehicle, to reduce to a minimum the risk of in-flight failures. Acceptance thermal vacuum could be cancelled after a period of consolidation of the manufacturing and AIT process, i.e. after the delivery of three to four flight units, together with the in-flight demonstration that the early flight failures are sensibly reduced. Acoustic tests are required for qualification and acceptance, while qualification sine vibration tests will be performed depending on the requirements of the selected launcher. Random vibration test applicability shall be evaluated following the same approach to be applied for the Hub. A full qualification model (QM) of the LEV active thermal control (ATC) and of the Environmental Control and Life Support (ECLS) subsystems is deemed necessary, to demonstrate the functions and performances of these subsystems, independent of the respective AV and Surface Module, and as fully integrated system on the LEV. Full EMC tests are also mandatory for the LEV. The most effective way to perform these tests, either at integrated LEV level or on its separated modules, may be defined in the next phases of the project, as part of the overall AIT planning.

10.4 Model philosophy A reference model philosophy is applied in the context of this study to provide a preliminary definition of which models of the system have to be built to fulfil the objectives of the test verification programme. The model philosophy is provided for the Hub and its submodules, and for the LEV and its elements. No model philosophy is defined here for the cargo vehicle or for the propulsion modules. Table 10-1 shows the following definitions:



page 283 of 317

s• Engineering Model (EM): a model that is functionally representative of the relevant flight

item (on-board active subsystems), and representative of its configuration, including harness and pipelines layout and sizing, habitation outfit etc. Redundancy is normally excluded.

• Structural Thermal Model (STM): a model representative (flight standard built) of structural and thermal characteristics of the relevant flight item, in which mechanical and thermal qualification tests are performed.

• Qualification Model: a model that is fully representative (flight standard) of the relevant flight item, down to components quality and reliability level. This model is used for full functional and performance qualification testing. It is also subjected to environmental testing and it will be evaluated if QM vehicles will be launched, to serve as first Flight Test Model (FTM).

• Flight Test Model (FTM): a full flight unit that is subjected to the complete cycle of integration and ground acceptance test activities, and is then launched to perform in-flight qualification of mission operations, including joint verification with the Mission Operation Centre (MOC).

Table 10-1: Model philosophy

10.5 Qualification flights A number of qualification flights shall be performed, in Earth orbit first, to verify the basic assembly or docking operations between modules and elements of the system. In the beginning, a partial configuration should be built up, progressively growing into the full system. All nominal and contingency flight manoeuvres shall be exercised and verified, including rendezvous, docking and separation. Flight vehicle on-board functions shall be verified and the actual system reliability checked out. All operations shall be exercised on board by the crew, or in remote mode by the MOC.

10.5.1 Hub

There should be qualification of flight operations in LEO: assembly operations between modules, RvD with the PMs, RvD with LEV, RvD with CTV, RvD with the cargo vehicle, and EVA operations. There should be qualification of transfer operations to LLO, modules docking and berthing, telecom in LLO.

10.5.2 LEV

There should be qualification of flight operations in LEO: assembly operations between modules, RvD with the PMs, RvD with Hub (full, and using the AV only), refuelling operations

Hub EM STM QM FTM1 FTM2 FTM3 FMModule 1 1 1 1 1 1 1Module 2 1 1 1 1 1 1

LEVDM 1 1 1 1 1 1 1Hab 1 1 1 1 1 1AV 1 1 1 1 1 1 1



page 284 of 317

s(if any, by the Cargo Vehicle), partial exploitation of descent operations. There should be qualification of transfer to LLO, telecom in LLO.

10.6 Ground training A number of ground simulators is foreseen and they have to be developed to train the crew and the ground support staff for this mission. Flight simulators shall be developed for the Hub and the LEV. A Moon Landing Trainer Vehicle shall be developed to allow the LEV pilots to practice with the very last operational phases of their descent to the Moon, until touch down. The vehicle shall be full scale and will provide adequate representation of flight properties and control capability. A Neutral Buoyancy Facility (NBF) shall be set up to support the system design first, and to verify and practice the manual operations later on. A training centre for the astronauts will be necessary. Lunar science including geological field research will be part of the instruction programme. CTV crew training is excluded in the context of this study.

10.7 Options Some manned flight operations that are planned for the lunar exploration mission could be verified on the ISS. For Hub docking operations with the LEV, a command module simulating the Hub could be attached to the ISS and LEV proximity operations (or just AV final docking operations) could be verified there.

10.8 Precursor missions A number of unmanned preliminary missions will be mandatory to satisfy the safety aspects and the effectiveness of the main lunar landing programme:

• Landing site candidates’ altimetry mapping, to improve the knowledge of the configuration of the terrain at the selected (or optional) landing sites. This is to reduce the risk on the LEV of touching down on a surface with critical slope or with distribution of craters exceeding the acceptable limits, or other geomorphic risks.

• Robotic rovers to investigate the landing sites.

10.9 Development schedule ID Task Name Duration2 Phase B 600 days3 Launch Manifest 1191 days26 Phase C/D 3130 days27 EM tests 230 days29 QM Tests 270 days31 FTM1 Flight Ops Verification 43 days34 FTM2 Flight Ops Verification 43 days37 HUB 2997 days38 Design 540 days39 HUB EM 810 days45 HUB QM 860 days51 HUB FTM1 740 days57 HUB FTM1 Launch and Flight Ops 421 days66 HUB FTM2 740 days72 HUB FTM2 Launch and Flight Ops 587 days89 LEV 3130 days90 Design 540 days91 LEV EM 810 days99 LEV QM 970 days107 LEV FTM1 1046 days116 LEV FTM 1 Launch and Flight Ops 340 days124 LEV FTM2 936 days131 LEV FTM2 Launch and Flight Ops 340 days139 LEV FTM3 1106 days148 LEV FTM3 Launch and Flight Ops 450 days

2 1 2 1 2 1 2 1 2 1 2 1 2 1 2 1 2 1 2 1 2 1 2 1 2 1 2 1 2 1 2 1 2 1-1 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16



page 285 of 317

sFigure 10-1: Overall project development schedule (up to last qualification flight)

The preliminary project schedule is based on the assumption that three missions with vehicle flight models are required to certify the capability of the system to fulfil the mission objectives and safely support manned flight. These models are called Flight Test Models (FTMs). The verification approach requires that the flight operations are first executed in automatic mode, and when they have been successfully accomplished, they are repeated on the next FTM by manual command of the on-board crew. This also includes joint operations with the CTV, but this part is not shown in the schedule. To allow the development of the system in a reasonable time, it is assumed that two different industrial consortia build the Hub and the LEV elements. A single core design team should ensure common interfaces. The time span of the project could be about 12 years development from the beginning of B phase until the launch of the first element of the Hub. The development activities continue in parallel with launches for another 4 to 5 years, during which the various flight modules and elements are delivered for launch, and mission operations are verified with the three planned FTMs of the Hub and LEV, together with the relevant propulsion modules. The launch manifest for the FTM elements of the system is shown here below. Cargo vehicle and CTV launches are not shown. Assuming some more risk (to be quantified), the LEV FTM2 manufacturing, ground and flight verification could be omitted, saving procurement budget and schedule time, provided that adequate effort is put into performance of the FTM1 test and operations. Another alternative is to use the LEV FTM3 and Hub FTM2 as FM1, performing with it final verification in automatic mode and then carrying the crew into LLO with the CTV, to perform the first full operational mission.

ID Task Name Duration3 Launch Manifest 1191 days4 Launch 1 HUB PM1 FTM1 0 days5 Launch 2 LEV DM FTM1 0 days6 Launch 3 HUB PM2 FTM1 0 days7 Launch 4 PropModule1 0 days8 Launch 5 LEV Hab/LAV FTM1 0 days9 Launch 6 PropModule2 0 days10 Launch 7 LEV DM FTM2 0 days11 Launch 8 LEV LAV FTM2 0 days12 Launch 9 PropModule3 0 days13 Launch HUB PM1 FTM2 0 days14 Launch Propulsion Module PM1-1 0 days15 Launch HUB PM2 FTM2 0 days16 Launch Propulsion Module PM1-2 0 days17 Launch Propulsion Module PM2-1 0 days18 Launch Propulsion Module PM2-2 0 days19 Launch LEV DM FTM3 0 days20 Launch Propulsion Module DM-FM3-1 0 days21 Launch Hab/LAV FTM3 0 days22 Launch Propulsion Module DM-FM3-2 0 days23 Launch Propulsion Module Hab/LAV FM3-1 0 days24 Propulsion Module Hab/LAV-FM3-2 Launch O 0 days25 Launch Hab/LAV FTM3 TLI/LLO 0 days

11-0801-1103-28

06-0608-0910-1212-15

05-2207-31

12-0402-1204-23

07-0210-25

01-1703-2305-2607-2910-01

01-1203-3006-02

2 1 2 1 2 1 2 1 2 1 2 1 2 1 2 1 2 1 2 1 2 1 2 1 2 1 2 1 2 1 2 1 2 1-1 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16

Figure 10-2: Launch manifest



page 286 of 317

s



page 287 of 317

s11 COST

This chapter describes the independent cost estimate for the lunar exploration mission. It includes cost estimates for Phase A, B and C/D for the following elements:

• Hub • Lunar Excursion Vehicle (LEV):

o Descent Module (DM) o Surface Habitation Module (SHM) o Ascent Vehicle (AV)

11.1 Class of estimate This cost estimate is classified, according to the ESA Cost Engineering Chart of Services, as Class 4 of a Major Complexity project, performed in a Normal time frame. This leads to design maturity margins that account for unknown design aspects not yet identified. These provisions are not risk margin (i.e. cost impacts due to the realisation of a stochastic event) and must be considered as part of the total industrial cost. Design maturity margins of 30% have been applied, due to a lack of available references.

11.2 Requirements and assumptions The scope of this cost estimate does not include all mission elements of the lunar exploration scenario. Elements such as launches, Crew Transfer Vehicle, Propulsion Modules (PM), Cargo Vehicle, international cooperation etc have not been taken into account. The resulting potential overall architecture and/or industrial environment is shown by the dotted lines in Figure 11-1. It was not in the scope of this analysis and might be studied at a later stage. It has been assumed that there will be separate “subprime” contractors for the four elements (DM, SHM, AV, Hub). Also will there be a prime responsible for the overall LEV project. The subprime contractors are assumed to be responsible for Assembly, Integration and Test (AIT) activity support at system level. A schematic set-up is shown in Figure 11-1.

Figure 11-1: Assumed industrial architecture



page 288 of 317

sThe cost estimates are based on a fully competitive environment with an optimised industrial architecture. Unless otherwise specified, the hardware will be based on off-the-shelf equipment or on existing and available technology as available at the time of the mission. Table 11-1 shows the reference model philosophy and the LEV elements as well as the Hub and its submodules. The selected model philosophy is based on Structural and Thermal Model (STM), Engineering Model (EM), Qualification Model (QM), Flight Test Model (FTM) and Flight Model (FM) for the assembly integration and verification process.

Table 11-1: Assumed hardware matrix

Note that there will be a total number of 14 LEVs and that the given launch sequence will force industry to produce in two parallel production lines. For the cost estimate, start of Phase B is foreseen for August 2007. Phase C/D is assumed to start in April 2009 and to end in October 1017. The study was performed for a sole ESA mission and therefore no international cooperation scenario has been taken into account. The design maturity cost margins of 30% account for additional costs resulting from later refinements of the design, revealing unseen complexities at the time of the estimate. To avoid piling up of several, different margins all input parameters for the cost models, such as mass, are used without margins. The cost estimates are based on economic conditions mid 2004. All cost estimates are based on references and cost estimating methods in line with the above general hypothesis. Geographical distribution effect is accounted for in the cost risk assessment.

11.3 Cost estimate methodology The following methods have been used, in descending order of preference:

• Reference to similar equipment/system level costs, taking into account the amount of new development required

• Bottom up analytical estimates for project teams • Expert judgement from technical specialists in combination with similar equipment

references, in the case that the amount of new development is extensive • Expert judgement from technical specialists only, if references are not available • Equipment cost models • Instrument cost models • The ESA internal, system level cost model RACE • System-level cost relationships (for the Prime and Subcontractor activities), based on

recently observed relationships for relevant references.



page 289 of 317

s11.4 Scope of the estimate The cost estimate includes costs for Hub and Lunar Excursion Vehicle:

• Phase A, B and C/D • Design maturity cost margins • Cost-risk margin • Geographical distribution constraints

The cost estimate does not include costs for:

• Consumables such as propellant, water, food etc • Any other Lunar Exploration Mission element such as a Crew Transfer Vehicle,

Propulsion Modules (PM), Cargo Vehicle, launches or international cooperation effects • Earth and lunar Ground Segment Operations and/or Launch and Early Orbit Phase

(LEOP) • ESA internal cost and contingencies

Table 11-2: Included and excluded mission elements in the cost estimate

11.5 Cost assessment For the cost estimate, each element is regarded as a project on its own handled by a subprime contractor at element system level. All Project Office (PO), AIT and Ground Support Equipment (GSE) costs are therefore included at element system level. Furthermore a LEV prime will have the same tasks and associated costs at overall LEV level.

11.5.1 Project Office

The Project Office costs at the different system levels include the costs for: • Management and Control (including overheads on subcontracts) • Product assurance • Engineering and documentation including payload interface engineering both at system

and subsystem level • Overheads on subcontractors, travel expenses

11.5.2 Assembly Integration and Verification, Ground Support Equipment

The AIV cost estimates include the costs for all system and subsystem mechanical and electrical integration activities and tests, as well as the mechanical mating of the different elements. The cost estimates are based on the specific subsystem and project team sizes, project durations, and yearly rates as well as CERs. The cost estimate for GSE covers the costs for all Electrical and Mechanical GSE required for the all elements. It has been taken into account that the GSE will be based on partly newly developed as well as mainly on existing hardware and designs. Accordingly, spacecraft GSE



page 290 of 317

scosts are based on discussions with experts and standard ratios observed in past projects. GSE maintenance has also been accounted for during the production phase of recurring models.

11.5.3 Hardware and software

The cost assessment has been characterised by a rather limited amount of available reference material for manned missions. Therefore whenever necessary and where unmanned references have been taken into account an adjustment factor has been applied based on discussions with experts and former estimate exercises such as a human mission to Mars. The design maturity cost margins have been influenced accordingly. Generally, costs for hardware and software are primarily based on the costs of prior ESA mission and internal CERs. Prices have been estimated based on this reference but are adjusted considering the programmatic constrains and with today’s market price trends. GNC equipment is assumed to be off-the-shelf with eventual simple modifications. Equipment analogy and ESA internal CERs have been used to estimate the costs. The propulsion subsystem development, production and test activities, are assumed to be performed by a propulsion subsystem sub-contractor. Equipment costs are based on TRANSCOST 7.1, FESTIP and prior ESA missions. To cope with the new structural configurations, costing has been based on the ESA internal cost model. Although solar cells for SHM and Hub will be off-the-shelf equipment, the panel configuration will be unique. The cost estimate therefore assumes that a normal solar array development effort is included. Battery, solar array and PCDU costs have been based on ESA internal CERs, all other elements on the Columbus power system. Harness costs have been determined using ESA internal CERs. TT&C costs have been estimated using analogy and ESA internal CERs. TT&C equipment is assumed to be off-the-shelf with eventual simple modifications. The CDMU has been assumed to be an existing and modified equipment unit. The camera costs are based on the cameras on board or planned for Rosetta or ExoMars. ESA internal cost models, CERs and analogy have been used to estimate the costs. The equipment cost estimates for the thermal subsystem are based on ESA internal CERs. Different equipment complexities have been taken into account. Mechanism costs are based on ESA internal CERs, where the necessary development effort has been taken into account. The Columbus ECLSS has been considered as the main reference for the life support system cost estimates. To derive the software cost estimates, the mission complexities have been considered. Main reference has been Ariane-5 software costs. A total number of 14 LEVs has been taken as reference. The two associated production lines will have two effects. Firstly, industry responsible for the LEV needs to have an accordingly larger team. Secondly, there will be learning effects from one team to another and with each production. The large industrial team has been taken into account in the core estimate. To cope with the learning effect, a standard learning curve derived from “Space Mission Analysis and Design” has been applied. Taking into account the first unit cost, the total number of 14 flight models (one non-recurring and 13 recurring to run the cost model correctly) and a typical learning curve factor of 0.95 leads to a cumulated total cost for DM, SHM and AV.



page 291 of 317

s The design maturity margins as described above account for unknown design aspects not yet identified. These provisions are not risk margins (i.e. cost impacts due to the realisation of a stochastic event) and must be considered as part of the total industrial cost. The design maturity margins are 30% for the LEV and 30% for the Hub.

11.6 Cost-risk analysis Full open competition hypothesis may not be retained, in particular for the prime and the first sub-contacting layer. Additional subcontracting layers with respect to cost efficiency may occur due to the magnitude of the investment. A 5% cost risk provision is foreseen for adjusting geo-return requirements at equipment level. Staggered phasing may occur triggered by slow decision making on programme funding. Planning disruption on the critical path at different stages of the project have been evaluated and implemented in the cost-risk analysis. For eventual launch failure during the time of the projects a provision has been foreseen. A probability of 5% and 2 years delay has been assumed. Not included are costs for the launcher. A provision for yet unidentified risks has been considered. The scope of this cost estimate does not include the overall mission architecture with all elements. Therefore the defined cost risk provision does not represent the full value to be considered for the whole programme.

Table 11-3: Cost-risk synthesis (all costs in %)



page 292 of 317

s



page 293 of 317

s12 CONCLUSIONS

Within this study two lunar exploration scenarios have been analysed and one baseline mission architecture selected that fits with the main objectives of both scenarios. The mission architecture has been studied in detail and two spacecraft elements have been designed. The other elements have been assessed at system level. The selected mission architecture features an orbital inhabited infrastructure (Hub) in lunar orbit (polar or equatorial depending on the scenario). The infrastructure acts as safe haven in case of a radiation storm for the crew on the surface of the Moon and allows testing of long-term habitation technologies and operations in deep space. A Lunar Excursion Vehicle, sized for a crew of three and for 14 days permanence on the surface at any location, performs the descent and landing and is used as surface habitation module. The architecture relies on a 27-tonne-into-LEO-class launcher that implies splitting the architecture vehicles into subelements, assembling the subelements to propulsion modules in LEO for transfer to lunar orbit and final assembly of the vehicles in lunar orbit. In the case of the sustainable exploration scenario, the selected mission architecture allows for more than 10 landings in 10 years and more than 1000 EVA hours on the surface. This mission architecture represents a step forward compared to the Apollo approach as regards the following aspects:

• Increased number of crew to the surface (three versus two) • Increased surface stay duration (14 days versus 2-3) • Presence throughout the whole mission of a safe haven (permanent infrastructure

orbiting around the Moon, Hub) allowing shelter within one day in the event of a radiation storm

As mentioned, a major difference (and significant drawback) is the assumption of unavailability of heavy lift launchers of the type of Saturn V. The performance parameter of the architecture is the landing rate because this quantifies the number of surface sites and number of surface days available for exploration. The landing rate is a function of the LEV assembly time (in LEO and in LLO) and in turn of the LEV mass: the lower the LEV mass, the higher the landing rate. However, the mass of the LEV cannot be reduced below a certain threshold driven by the requirements of 14 days surface permanence for three crewmembers. This study has shown how this parameter can be optimised. The study identified the following main issues of the selected architecture:

1. The requirement of the capability of landing anywhere on the Moon is a major driver. This restricts the choice for the location of the Hub to either a LLO polar orbit (preferred option in this study) or the L1 point between Earth and Moon. In the first case, landing windows exist which reduce the landing rate capability of the architecture and there is an additional ∆V penalty for orbit maintenance. In the second case, large ∆V for descent to surface is needed, increasing the mass of LEV.



page 294 of 317

s2. The limitation of launch capability leads to a high number of launches and complex

assembly operations. The availability of a launcher in the 80-tonne-into-LEO-class would greatly simplify the architecture.

12.1 Technology development Conservative assumptions on technology availability and performance have been made throughout the whole exercise. That is, innovation was not taken as a programme requirement. Even so, major technology developments are needed to achieve the mission. Among those, some enabling ones are:

• Closure of the life support subsystem • Techniques for reduction of boil-off in cryogenic propulsion system • High-thrust, high-throattability descent and landing engines • Automatic assembly techniques • Fuel cells • Advanced avionic systems and architectures

12.2 Complementary infrastructure Within the architecture proposed in this study there is a need for two robotic infrastructures:

1. A set of data relay satellites to guarantee communications with the far side of the Moon 2. A space-weather monitoring system consisting of several spacecraft and a grand

infrastructure in charge of monitoring solar activity and the Earth’s magnetic field and to warn of any possible impending radiation storm



page 295 of 317

s

13 REFERENCES

RD[1] Lo, M.W., Ross, S.D.: The Lunar L1 gateway: portal to the Stars and Beyond, AIAA Space Conference, Albuquerque, August 2001

RD[2] Condon, G.L., Wilson, S., Ranieri, C.L.: Performance Requirements and Kickstage Disposal Options for a Cislunar Gateway Station Vehicle, NASA-JSC

RD[3] Konopliv, A.S., Asmar, S.W., Carranza, E., Sjogren, W.L., Yuan, D.N.: Recent Gravity Models as a Result of the Lunar Prospector Mission, Submitted to Icarus, 1999

RD[4] Meyer, K.W., Buglia, J.J., Desai, P.N.: Lifetimes of Lunar Satellite Orbits, NASA Technical paper 3394, NASA Langley, March 1994

RD[5] Hoffman, D.A.: Station-Keeping at the Collinear Points of the Earth-Moon-System, NASA-JSC-26189, September 1993

RD[6] Gomez, G. et al.: Study Refinement of Semi-Analytical Halo Orbit Theory, Study report under contract to the European Space Agency, ESOC Contract Number: 8625/89/D/MD(SC), Technical Supervisor J. Rodriguez-Canabal, Aril 1991

RD[7] Bennett, F.V.: Apollo Lunar Descent and Ascent Trajectories, NASA TM-X-58040, March 1970

RD[8] Guidelines & Capabilities for Designing Human Missions, NASA/TM-2003-210785

RD[9] Columbus-CC & MSM Communications Infrastructure ADD, OPS-ADD-1CC001-GSOC, Issue 4 Upd.1, 16 February 2004

RD[10] ATV Rendezvous Pre-development Program, www.estec.esa.nl/wawww/ESC/arp.html

RD[11] NASA TN D-6846 Apollo Experience Report – Mission Planning for Lunar Module Descent and Ascent, by Floyd V. Bennett – June 1972

RD[12] NASA Project Apollo Working Paper No 1084 – Study of powered Descent Trajectories for Manned Lunar Landings – August 1963

RD[13] Details on Apollo Spacecraft- document.

RD[14] http://www.fourmilab.ch/earthview/lunarform/cratall.html, Named Crater Details [1].

RD[15] http://planetarynames.wr.usgs.gov/ , Lunar Named Crater Details [2].

RD[16] http://www.enchantedlearning.com/subjects/astronomy/moon/Craters.shtml , Crater Geometry Detail.

RD[17] ESA PSS-03-406 Issue 1 (august 1991) Life Support and Habitability Manual. Volume 1. Section 1.4, pages 6-7

RD[18] MELiSSA Technical Note 14.1 Simulation des flux de matiere dans l’écosystéme MELiSSA, page 4



page 296 of 317

sRD[19] Apollo Experience Report. Consumables budgeting. NASA TN-D-7140

RD[20] Foundations of Space Biology and Medicine Volume III (NASA SP-374) Page 19

RD[21] The Mars Project: Avoiding Decompression Sickness on a Distant Planet. NASA TM-2000-210188,

RD[22] Baseline Values and Assumptions Document, NASA CD-2004-208941 (August 2004)

RD[23] www.hsssi.com (last visit, December 15th)

RD[24] Living together in Space: The Design and Operation of the Life Supprt Systems on the International Space Station, NASA TM-1998-206956, Volume 1

RD[25] Baseline Values and Assumptions Document, NASA CD-2004-208941 (August 2004)

RD[26] www.hsssi.com (last visit, December 15th)

RD[27] Living together in Space: The Design and Operation of the Life Support Systems on the International Space References

RD[28] Callot. S., “Aurora-Mars Sample Return / Assessment of chemical propulsion system for interplanetary missions ”, TN4-WP3100, Snecma Moteurs, 2004

RD[29] Cecil. R. & Wood. J., “Apollo Experience Report-Service Propulsion Subsystem”, TN D-7375, NASA, 1973

RD[30] Bennet. F., “Apollo Experience Report-Mission Planning for Lunar Module Descent and Ascent”, TN D-6846, NASA, 1972

RD[31] Hooper. J., “Performance Analysis of the Ascent Propulsion System of the Apollo Spacecraft”, TN D-7400, NASA, 1973

RD[32] Hammock. W. & Currie. E., “Apollo Experience Report-Descent Propulsion System”, TN D-7143, NASA, 1973

RD[33] Huzel. D. & Huang. D., “Modern Engineering for Design of Liquid-propellant Rocket Engines”. Rocketdyne Division of Rockwell International

RD[34] Rocketdyne Propulsion & Power, “Liquid Propellant Rocket Propulsion System”. Boeing

RD[35] Education Department EXR/E., “Man Mission to the Moon”, System Description Document, M3-PA-SDD-0707-1-25, ESA/ESTEC.

RD[36] Wojtowicz, J.A. “Dynamic Properties of the International Space Station throughout the Assembly Process,” 98-027, Jun. 98

RD[37] Analysis and Design of Missile Structures, E.F. Bruhn

RD[38] ECSS-E-30 Part 2A, Space Engineering, Mechanical-Part 2: Structural

RD[39] Columbus System Requirement Document

RD[40] Human Mission to Mars CDF Study Report

RD[41] Seminarabeit Transhab Inflatable Habitat



page 297 of 317

sRD[42] Plancke P., David P., Technical dossier “On Board Computer & Data Systems”.

European Space Technology Harmonisation. ESA.

RD[43] “Human Mission to Mars”. CDF Study. ESA.

RD[44] Pouponnot, A. ESA Microprocessors Roadmap. ESA.

RD[45] CCSDS SOIS Website. http://www.ccsds.org

RD[46] “Guidelines and Capabilities for Designing Human Missions”. NASA/TM-2003-210785

RD[47] Interplanetary Internet http://www.ipnsig.org

RD[48] Wearable computer. http://www.xybernaut.com/Solutions/product/listing_product.htm

RD[49] Wireless in space http://www.wireless.esa.int

RD[50] Spacecraft Solar Array Technology Trends, P. Alan Jones & Brian R. Spence, AEC-Able Engineering Company, Inc. Goleta

RD[51] Apollo Experience Report: The Cryogenic Storage System, NASA Technical Note, June 1973 W.A. Chandler, R.R. Rice, R.K. Allgeier Jr, NASA, Houston

RD[52] PM1: Structural Design of Advanced Solar Array, Contract ESA 16714/03/NL/CH, 6th November 2003 Alcatel Space

RD[53] Thin Film Solar cells – Phase A: Flexible solar array concepts, ASPI-02-RT/PS-25, 9th July 2002 M. Tur, Alcatel Space

RD[54] Advanced Photovoltaic Solar Array Program Update, JPL/NASA/TRW, 23-27 August 1993 R. Kurland (TRW) and P. Stella (JPL), 3rd ESPC 93 Graz, Austria

RD[55] The Lunar Base handbook, ISBN 0-07-240171-0 Peter Eckart, Space Technology Series

RD[56] Design Considerations for Lunar Base Photovoltaic Power Systems J.M. Hickman, H.B. Curtis, G.A. Landis, NASA Lewis Research Center

RD[57] Lunar Sourcebook: A User’s Guide to the Moon G. Heiken, D. Vaniman, B.M. French, Cambridge University Press

RD[58] Lunar Lander/Rover: Power Subsystem: Concept Definition Study, EWP-1843, June 1995 J.H.I. Boks, ESTEC Working Paper, ESA Power & Energy Conversion Division

RD[59] The Fuel Cells in Space: Yesterday, Today and Tomorrow, NASA Technical Memorandum 102366, 18-21st September 1989 M. Warshay and P.R. Prokopius, NASA Lewis Research Center

RD[60] CDF Study Report: Human Missions To Mars: Overall Architecture Assessment, CDF-20(A), February 2004, CDF Team, ESTEC/ESA

RD[61] CDF Study Report: Human Spaceflight Vision: European Man-tended Moon Base, CDF-23(A), January 2004, CDF Team, ESTEC/ESA



page 298 of 317

sRD[62] Non Nuclear Energy Storage Systems for Various Applications On Mars or Others

Planetary Missions, 6th March 2001 Max Schautz, ESA, TEC-EPB

RD[63] The Corrugated thin film solar cell design, 27-28th February 2003 R. Zwanenburg, Dutch Space, Thin Film Solar Arrays Workshop, ESA, Noordwijk

RD[64] Future Power Systems for Space Exploration: Executive Summary, February 2002 Adam M. Baker, Qinetiq

RD[65] Future Power Systems for Space Exploration Technical Note 2: Technology Inventory, November 2001, QINETIQ/KIS/SPACE/TN011164 Adam M. Baker, Qinetiq

RD[66] Future Power Systems for Space Exploration Technical Note 4: Architectural Design of Mars Surface Power System, February 2002, QINETIQ/KIS/SPACE/TN010446

RD[67] Flywheels Ready To Fly Spacedaily, 10 April 2000

RD[68] Cryogenic Reactant Storage For Lunar Base Regenerative Fuel Cells Lisa L. Kohout, Nasa Technical Memorandum 101980

RD[69] ECSS-M-00-03A Risk management

RD[70] ECSS-Q-40 Safety

RD[71] NASA/TM-2003-210785 Human Rating requirements.

RD[72] SF24-53/I. Selection of Spectrum Candidates in the Vicinity of the Moon. SFCG.

RD[73] ‘Nasa/Gsfc Ground Segment Upgrades For Ka-Band. Support To Near-Earth Spacecraft’. Yem Wong, Mark Burns. SpaceOps2002 Papers. http://www.spaceops2002.org/papers.html

RD[74] ITU Radio Regulations. ITU RR5. Frequency allocations.

RD[75] ‘Radio frequency and modulation systems, Part.1: Earth Stations and Spacecraft’. Recommendation CCSDS 401x0b11.

RD[76] ECSS-E-50 Part 1A, Communications — Part1: Principles And Requirements, published 20 October 2003

RD[77] ECSS-E-50-05A, Radio Frequency And Modulation Standard, published 24 Jan 2003

RD[78] CCSDS 101.0-B-6. Telemetry Channel Coding Standard. Blue Book. Issue 6. October 2002.

RD[79] PSS-04-104, Ranging Standard, March 1991. It will be superseded with ECSS-E-50-02 standard.

RD[80] CCSDS 102.0-B-5. Packet Telemetry Standard. Blue Book. Issue 5. November 2000.

RD[81] CCSDS 201.0-B-3. Telecommand Part 1—Channel Service. Blue Book. Issue 3. June 2000.



page 299 of 317

sRD[82] CCSDS 202.0-B-3. Telecommand Part 2—Data Routing Service. Blue Book. Issue

3. June 2001.

RD[83] CCSDS 202.1-B-2. Telecommand Part 2.1—Command Operation Procedures. Blue Book. Issue 2. June 2001.

RD[84] CCSDS 203.0-B-2. Telecommand Part 3—Data Management Service. Blue Book. Issue 2. June 2001.

RD[85] ECSS-E-70-41A, Telemetry And Telecommand Packet Utilization, published 30 January 2003

RD[86] CCSDS 211.0-B-2. Proximity-1 Space Link Protocol—Data Link Layer. Blue Book. Issue 2. April 2003.

RD[87] CCSDS 211.1-B-1. Proximity-1 Space Link Protocol—Physical Layer. Blue Book. Issue 1. April 2003

RD[88] CCSDS 211.2-B-1. Proximity-1 Space Link Protocol—Coding and Synchronization Sublayer. Blue Book. Issue 1. April 2003

RD[89] The Temporal and Spectral Characteristics of Ultrawideband Signals. S.Department Of Commerce. NTIA Report 01-383

RD[90] Simon Fraser. Communication Systems Architectures and Analog Communication Systems Architectures and Analog Mission Testing for Advanced Robotic and Human Mission Testing for Advanced Robotic and Human Missions to the Missions to the Moon. ESA academia lunar workshop 8-9 november 2004

RD[91] TDRSS Technical Info Package, NASA. http://nmsp.gsfc.nasa.gov/TUBE/techinfo.htm

RD[92] Aurora Communications Roadmap. Jean-Luc Gerner. 4 February 2004. Technical Note, Aurora Communications Team.

RD[93] Performance Comparison of Different bandwidth-Efficient Modulation Schemes for Earth Exploration Satellite TM Systems at 8 GHz. M. Martinez Fernandez. ESA-ESTEC, TTC and Radionavigation Section (D/TEC-ETT). TT&C workshop 2004.

RD[94] Cisco Aeronet Omnidirectional mast Mount antenna (AIR-ANT2506). http://www.cisco.com/univercd/cc/td/doc/product/wireless/acessory/ant2506.pdf



page 300 of 317

s



page 301 of 317

s14 ACRONYMS

AIT Assembly Integration and Test AM Ascent Module AM0 Air Mass Zero ANITA ANalysing InTerferometer for Ambient air AOC Attitude and Orbit Control AOCS Attitude and Orbit Control System APS Ascent Propulsion System ARES Air REvitalisation System ATC Active Thermal Control ATC Astronaut Training Centre ATM Asynchronous Transfer Mode ATV Automatic Transfer Vehicle BCR Battery Charge Regulator BDR Battery Discharge Regulator BOL Beginning of Life CAN Controller Area Network CC Control Centre CDMU Control and Data Management Unit C-I/F Crew interfaces CoG Centre of Gravity CoM Centre of mass COPV Composite Over-wrapped Pressure Vessels COT Crew Operations Team COTS Commercial Off The Shelf CRV Crew Return Vehicle CTV Crew Transfer Vehicle CV Cargo Vehicle DGPS Differential GPS DHS Data Handling System DLS Descent and Landing System DM Descent Module DoD Depth Of Discharge DOI Descent orbit injection DOR Differential On-way Ranging DPS Descent Propulsion System DSM Deep Space Manoeuvre EADS European Aeronautic Defense and Space Company



page 302 of 317

sEAM European Apogee Motor ECLS Environmental Control and Life Support EDAC Error Detection and Correction EEPROM Electrically Erasable Programmable ROM EL-CC Earth-Lunar Control Centre EM Engineering Model EMS ESTRACK Management System EOL End Of Life ESC Engineering Support Centre ESTRACK ESA Tracking Network ETO Earth Transfer Orbit EVA Extra Vehicular Activity FC Fuel Cell FCT Flight Control Team FDIR Failure, Detection, Isolation and Recovery FDS Flight Dynamics Support Fh Resulting Horizontal Force at Landing FM Flight Model FPA Flight Path Angle FRT Free Return Trajectory FS Factor of Safety FT Fault Tolerance FTM Flight Test Model Fv Resulting Vertical Force at Landing GaAs Gallium Arsenic GPS Global Positioning System GS Ground station HEPA High Efficiency Particulate Air filter HGA High Gain Antenna HK House-Keeping HKTM Housekeeping Telemetry HM Habitation Module IGS Interconnection Ground Subnet IMU Inertial measurement unit IMV Inter Module Ventilation iRED Interim Resistive Exercise Device ISDN Integrated Services Digital Network ISL Interspacecraft link ISS International Space Station



page 303 of 317

sL1 Lagrange Point 1 LAN Local Area Network LBNP Lower Body Negative Pressure LEO Low Earth Orbit LEOP Launch and Early Operations Phase LEV Lunar Excursion Vehicle LGA Low Gain Antenna LLO Low Lunar Orbit LM (Apollo) Lunar Module LOI Lunar Orbit Insertion LS-CC Lunar Surface Control Centre LTO Lunar Transfer Orbit MA Multiple Access MCA Major Constituents Analyser MCS Mission Control System MDPS Meteoroid and Debris Protection System MEOP Maximum Expected operating Pressure MIDDASS Microbial Detection and Sampling System MLI Multi Layer Insulation MLTV Moon Landing Trainer Vehicle MMDR Multimode Doppler radar MMH Mono Methyl Hydrazine MMT Mission Management Team MOC Mission Operation Centre MOI Moment of inertia MON Mixed Oxidisers of Nitrogen MPPT Maximum Power Point Tracker MSM Manned Spaceflight and Microgravity MSM-CI MSM-Communications Infrastructure MTL Mission TimeLine NRT Near Real Time OBA On Board Autonomy OBC On-Board Computer OBSW On-Board Software OFDM Orthogonal Frequency Division Multiplexing PDI Powered Descent Initiation PEMFC Proton Exchange Membrane Fuel Cell PFM Protoflight Model PLSS Portable Life Support System



page 304 of 317

sPM Propulsion Module PMD Propellant Management Device PVA Photovoltaic Array QM Qualification Model RCS Reaction Control System RCT Reaction Control Thruster RDM Rendezvous Manoeuvre RFC Regenerative Fuel Cell RFN Radiofrequency navigation RR Retroreflector RS Reed-Solomon RSOFC Regenerative Solid Oxid Fuel Cell RV Refuel Vehicle RVD Rendezvous and docking RVS Rendezvous sensor S/A Subassembly S/C Spacecraft S3R Sequential Switching Shunt Regulator S4R Sequential Switching Shunt Series Regulator SA Solar Array SA Single Access SCT Science Co-ordination Team SHM Surface Habitation Module SLE Space Link Extension SOFC Solid Oxid Fuel Cell SPF Single Point Failure SSO Sun Synchronous Orbit STM Structural Thermal Model TBD To Be Determined TC Telecommand TDRSS Tracking and Data Relay Satellite System TEI TransEarth Injection TJ Triple Junctions TLI Translunar injection TT&C Telemetry, Tracking & Command UPS Unified Propulsion System USOC User Support Operations Centre UWB Ultra Wide Bandwidth Vh Horizontal Velocity in m/s



page 305 of 317

sVPN Virtual Private Network Vv Vertical Velocity in m/s w.r.t. With respect to WMAN Wireless Metropolitan Area Networks XPDR Transponder



page 306 of 317

s



page 307 of 317

s

APPENDIX A: Lander stability model

Lander Static Stability Figure A-1 defines the geometry for the static stability for the N-leg landing system, assuming that the spacecraft lands with at least one leg on a rock of a given height as the worst case:

r

R’

H

h

x1

i

CoM

g mars

Θ

ρ

β

Figure A-1: Geometry for static stability analysis

The minimum footprint radius ‘r’, without margin, required to maintain stability when at rest is given by the following equation:

( )

⎟⎟⎟⎟⎟

⎠

⎞

⎜⎜⎜⎜⎜

⎝

⎛

⎥⎥⎥⎥

⎦

⎤

⎢⎢⎢⎢

⎣

⎡

⎟⎟⎠

⎞⎜⎜⎝

⎛⎥⎦⎤

⎢⎣⎡+

+

⎥⎦⎤

⎢⎣⎡+

=⇒

Nr

Hi

N

xhr

.2360cos1

arcsintan

.2360cos

1

The final angle w.r.t. the Martian surface, at which the lander will come to rest is given by the following equation:

⎥⎥⎥⎥

⎦

⎤

⎢⎢⎢⎢

⎣

⎡

⎟⎟⎠

⎞⎜⎜⎝

⎛⎟⎠⎞

⎜⎝⎛+

=Θ

Nr

H

.2360cos1

arcsin



page 308 of 317

sFor the four leg landing system and any configuration with an even number of legs, two situations can be considered:

• One side (two feet) land upon a rock • One foot lands upon a rock and the diametrically opposite leg contacts the ground

For configuration 1:

R’

H

h

x1

i

CoM

g mars

Θ

ρ

R’

Figure A-2: Landing on one foot configuration (side)

The minimum footprint radius ‘r’, without margin, required to maintain stability when at rest is given by the following equation:

( )

⎟⎟⎟⎟

⎠

⎞

⎜⎜⎜⎜

⎝

⎛

⎥⎥⎥⎥

⎦

⎤

⎢⎢⎢⎢

⎣

⎡

⎥⎦⎤

⎢⎣⎡

+

⎥⎦⎤

⎢⎣⎡+

=⇒

Nr

Hi

N

xhr

.2360cos..2

arcsintan

.2360cos

1

For configuration 2:

( ) ⎟⎟⎠

⎞⎜⎜⎝

⎛⎥⎦⎤

⎢⎣⎡++=⇒

rHixhr.2

arcsintan1

The final angle w.r.t. the Martian surface, at which the lander will come to rest is given by the following equation:

⎟⎟⎟⎟

⎠

⎞

⎜⎜⎜⎜

⎝

⎛

⎟⎠⎞

⎜⎝⎛

=Θ

Nr

H

.2360cos..2

arcsin and ⎟⎠⎞

⎜⎝⎛=Θ

rH.2

arcsin



page 309 of 317

s The static stability of this configuration is further complicated by the fact that the lander can topple or rotate around the axis created by the foot on the rock and the foot contacting the ground. The result will be a three-leg contact with the surface, with the remaining leg suspended off the surface. For a given footprint dimension, the topple angle can be determined as follows:

Ψ H/2

h+x1

R”

x

R”

β

Figure A-3: Footprint dimensions calculation

From the above geometry, the rotation Ψ is:

⎟⎟⎟⎟

⎠

⎞

⎜⎜⎜⎜

⎝

⎛

⎟⎠⎞

⎜⎝⎛

=Ψ

Nr

H360sin2

arcsin and ⎟⎠⎞

⎜⎝⎛=Ψ

rH.2

arcsin

If there is a slope under the lander, the lander will topple further until the foot contacts the surface. The above equation needs to be modified by an additional variable as follows:



page 310 of 317

s

Ψ

R”

H/2

h+x1

hi i

x

Figure A-4: Topple angle

From the above geometry, the rotation Ψ is:

⎟⎟⎟⎟

⎠

⎞

⎜⎜⎜⎜

⎝

⎛

⎟⎠⎞

⎜⎝⎛

+⎟⎠⎞

⎜⎝⎛

=Ψ

Nr

hHi

360sin

2arcsin and

⎥⎥⎥⎥

⎦

⎤

⎢⎢⎢⎢

⎣

⎡+⎟

⎠⎞

⎜⎝⎛

=Ψr

hHi2arcsin

The term hi is given by:

( ) ⎥⎥⎦

⎤

⎢⎢⎣

⎡⎟⎟⎠

⎞⎜⎜⎝

⎛+−⎟

⎠⎞

⎜⎝⎛−⎟⎟

⎠

⎞⎜⎜⎝

⎛⎟⎠⎞

⎜⎝⎛=

ihH

Nr

Hh ii 2

222

tan11

2360sin1 and ( ) ⎥

⎥⎦

⎤

⎢⎢⎣

⎡⎟⎟⎠

⎞⎜⎜⎝

⎛+−⎟

⎠⎞

⎜⎝⎛−=

ihHr

Hh ii 2

22

2

tan11

21

Lander Dynamic Stability Assumption 1- Vertical velocity, Vv = 0 Figure A-5 shows the geometry for the dynamic stability for the N-leg landing system:



page 311 of 317

s

Vh

x1

h

R’ r

contact

Ω = 0 rads/s

Stability limit with zero KE

HH’

gmars

Figure A-5: Dynamic stability definition, horizontal velocity case

The following analysis assumes the lander is travelling as shown in Figure A-6 and is therefore the worst-case (both or one-foot contact and no rotation about the contact foot assumed).

R

R’

β

Figure A-6: Landing with lateral velocity (3 legs case)

The following equations result in:

( ) ( )21

2

1

2

.21

2360cos

1 xhxhgVh

N

rmars

+−⎥⎦

⎤⎢⎣

⎡++

⎟⎠⎞

⎜⎝⎛

=

For the four-leg landing system and configurations with an even number of legs, two configurations can be considered:

• One side (two feet) contact the ground • One foot contacts the ground



page 312 of 317

s

R’

r

1

2

Figure A-7: Landing with lateral velocity (four legs case)

For configuration 1, the stability is given by the same model valid for the three-leg system. For configuration 2:

( ) ( )21

2

1

2

.21 xhxh

gVhr

mars

+−⎥⎦

⎤⎢⎣

⎡++=

Assumption 2- Vertical velocity, Vv = 0, Lander has attitude error For the N-leg landing system, the dynamic equation becomes: (H’’ is the height of the CoM including the height increase due to the attitude error)

( )( ) ( )2

1

2

1

2

cos.21

sin12360cos

1 xhxhgVh

N

r Amars

A

+−⎥⎦

⎤⎢⎣

⎡Θ++

Θ−⎟⎠⎞

⎜⎝⎛

=

For the four-leg landing system, two configurations can be considered:

• One side (two feet) contact the surface • One foot contacts the surface

For configuration 1, the stability is given by the above model. For configuration 2:

( ) ( ) ( )21

2

1

2

cos.21

sin11 xhxh

gVhr A

marsA

+−⎥⎦

⎤⎢⎣

⎡Θ++

Θ−=

Assumption 3- Vertical velocity, Vv > 0 Figure A-8 shows the geometry for the dynamic stability of the N-leg landing system:



page 313 of 317

s

Vh

x1

h

R’ r

contact

Ω = 0 rads/s

Stability limit with zero KE

HH’

gmars Vv

Figure A-8: Dynamic stability definition, horizontal and vertical velocity case

The following equation results:

( ) ( )21

2

1

22

.21

2360cos

1 xhxhg

VvVh

N

rmars

+−⎥⎦

⎤⎢⎣

⎡++

−

⎟⎠⎞

⎜⎝⎛

=


• One side (two feet) contact the surface • One foot contact the surface

For configuration 1: the stability is given by the above model. For configuration 2:

( ) ( )21

2

1

22

.21 xhxh

gVvVhr

mars

+−⎥⎦

⎤⎢⎣

⎡++

−=

Assumption 4- Vertical velocity, Vv > 0, Lander has attitude error The following equations are stated by inspection of the above equation derivation; For the N-Leg lander configuration:

( )( ) ( )2

1

2

1

22

cos.21

sin12360cos

2 xhxhg

VvVh

N

r Amars

A

+−⎥⎦

⎤⎢⎣

⎡Θ++

−

Θ−⎟⎠⎞

⎜⎝⎛

=


• One side (two feet) contacts the surface



page 314 of 317

s• One foot contacts the surface

For configuration 1, the stability is given by the above model. For configuration 2:

( ) ( ) ( )21

2

1

22

cos.21

sin11 xhxh

gVvVhr A

marsA

+−⎥⎦

⎤⎢⎣

⎡Θ++

−Θ−

=

Load on Lander leg Additionally, the model has been expanded to estimate the maximum load on the leg system for the given landing parameters assuming that all the landing load is taken on a single leg. The following approach has been taken to estimate the maximum load on the leg due to landing.

CoG

Fh

α

Mass, Vv, gLunar

Fleg.Axial

Fleg.Bending

Fv

Vh

Figure A-9: Loads calculation sketch

The following equations are relevant to calculate the force on the leg:

( ) ( )( ) ( )αα

αα

cossin

sincos

*

*

.

.

FhFvF

FhFvFa

VhmassF

ga

VvmassFv

Bendingleg

Axialleg

H

Lunar

+=

+=

⎟⎠⎞

⎜⎝⎛=

⎥⎦

⎤⎢⎣

⎡+⎟

⎠⎞

⎜⎝⎛=

The following assumptions are made:

• Assume deceleration time the same for both velocity components. • This approach excludes any lander attitude errors • This method assumes that the leg is placed along the line connecting the surface

contact point with the lander’s CoG point This method assumes no damping in the landing system.



page 315 of 317

s

APPENDIX B: Mission architecture model

Within this study model has been developed to carry out mission architecture trade-offs. The model has been called: super model. It allows performing trade-offs at architecture level in the early stages of the study as well as performing the integration of all the different mission elements data during and at the end of the study. The super model is based on Excel, for compatibility with the rest of the CDF Integrated Design Model. It is enclosed in a single workbook, with links to the dedicated models of each vehicle composing the architecture. The structure of the model is shown in Figure B-1:

Mission Analysis database

Architecture definition

Launchers database

Vehicles database and models

Mission definition

Scenario options

definition

Architecture Database Scenario

options trade offs

Figure B-1: Super model structure

• Input data (in grey): o Mission analysis database, containing all the ∆V and trip durations required to

design the mission, as provided by mission analysis specialists o Launcher database, containing the main performances of the launchers selected for

the study (mass to LEO, mass into LTO, launch rate, etc.) o Vehicle database, containing the major characteristics of past missions vehicles (i.e.

Apollo) and simplified models for sizing the different mission vehicles as a function of the number of crew, the mission duration and the ∆V.

• Architecture definition, where parameters such as number of mission elements (missions required), number of crew, surface/orbit durations and trajectories can be defined. A first



page 316 of 317

sestimation of the total mass, assembly time and number of launches required and some other user-defined parameters can be obtained. The results computed are stored in the Mission architecture database for later trade-offs.

InputsMission objectives f

Storable 0.08Cryo 0.11

ParameterRequirements Value UnitMinimum number of crew 3 [#]Minimum number of crew landed 2 [#]Minimum surface duration 14 [days]Landing date 1/6/2020Landing site latitude 0 [deg]Landing site longitude 0 [deg]Maximum mission duration 200 [days]Mobility on surface 0 [km]

Mission elementNameSubelement nameSubelement nameSubelement nameSubelement nameSubelement nameDefinitionCrew landed 3Number of crew [#] 6

Misson analysis selection Via L1? 2 Via L1?

Direc to lunar surface? NO Direc to lunar surface?

Direct from lunar surface? YESDirect from lunar surface?

Lunar orbit altitude [km] 4

Earth to Moon 4 Direct launch from Equator 1

Earth to L1 4 Direct launch from Equator 2

L1 to Moon 7 Not going 7

Moon to Earth 7 Not coming back 7

Moon to L1 7 Not coming back 7

L1 to Earth 7 Not coming back 7

Lunar ExcHub

Apollo validation

EleElement 1

Option 4

Option 4

Option 7

Option 7

Option 7

Option 7

100

NO

Option 1

Option 2

Option 7

Option 7

Option 7

Option 7 Figure B-2: Mission architecture worksheet screenshot

• Mission definition, for each of the mission included in an architecture, more detailed calculations can be performed with this model, that includes, for example, the gravity losses and boil-off analysis.

The results for the different options (stored in the Architecture database) can be compared using the Architecture options trade-off sheet, which takes as input the architecture database and the option definition sheet, in which architecture options and trade-off criteria can be defined. This sheet allows the user to select and use different trade-off criteria and weighting systems.



page 317 of 317

s

Figure B-3: Trade-off table

The super model is a dynamic tool for a first assessment of different mission architectures for complex missions. Once an architecture is selected and the relevant vehicle design is performed in more detail within the dedicated CDF model, the trade-off can be updated.

Parameter Units Definition Weight mark if threshold CHOOSE Value Mark Value Mark

1 > 2300002 = 2300003 < 2300001 > 152 ⊂ 5 to 153 < 51 > 3002 = 3003 < 3001 > 102 ⊂ 5 to 103 < 51 < 0.802 = 0.803 > 0.801 < 0.302 = 0.303 > 0.301 < 1502 = 1503 > 1501 < 1502 = 1503 > 1501 < 1502 = 1503 > 1501 < 1502 = 1503 > 150

10

TOTAL 56 50

MIN mark to MAX value

MIN mark to MIN value

MIN mark to MIN value

THRESHOLD

THRESHOLD

THRESHOLD

B2.00

569100 2.4

4 3.0

1 3.0

1 0.3

25 3.0

1.00 0.3

1.43 3.0

1 0.3

A1.00Marking

MIN mark to MAX value

Total launched mass/NOF [kg/#] 10

Number of objectives fulfilled (NOF) [#] 10

4 3.0

Total number of Launches/NOF [#/#] 1

1334890 3.0

53 3.0

1.43 3.0

Number of mission elements/NOF [#/#]

Ratio operational time/total time

[days/days] 5

1

Technology maturity by mission date [#] 5

Safety [?] 0

1.00 0.3

1 3.0

None [#] 0 THRESHOLD 0 1.0 0 1.0

None [#] 0 MIN mark to MAX value 0 0.3 0 0.3

None [#] 0 MIN mark to MAX value 0 0.3 0 0.3

#REF! #REF!

cdf study report - european space agencyemits.esa.int/emits-doc/1-5200-rd19-les_total.pdfles cdf...

Documents