human mars missions

Human Mars Missions

Paul Wooster

[email protected]

July 18, 2007

Getting to and from Mars

Earth-Mars-Earth Transportation Pathways

Earth Surface

Earth Orbit

Earth-Mars Transfer Mars-Earth Transfer

Mars Orbit

Mars Surface

Likely primary mission pathway

Unlikely pathway Potential abort pathway

Direct Return

The crew uses the TSH for Earth-Mars transfer and Mars surface stay

Return from Mars surface to Earth in ERV Only feasible if ISPP is utilized because of

large ERV habitat mass Example: Zubrin’s “Mars Direct”

2 vehicle designs required: TSH, ERV Mars surface rendezvous, all assets on the

surface and accessible to crew Split mission with pre-deployment is

required because of the use of ISPP: ERV needs to be fully fueled and ready for

Earth return before crew leaves Earth

Marssurface

Marsorbit

ERV TSH

Required crew transferbetween vehicles

Broken lines: uncrewed operations

Solid lines: crewed operations

Different colors indicatedifferent vehicles

Direct Return + Surface Habitat

Earth-Mars transfer in Transfer Habitat 1

Surface stay in surface habitat (SH)

Return from Mars surface to Earth in transfer habitat (TH-2) deployed one opportunity before

This strategy is only feasible with ISPP because of large return habitat mass

Mission mode requires either a single landing site or a chain of landing sites

2 vehicle designs required: SH and TH

Mars surface rendezvous, all assets on the surface and accessible to crew

Split mission required due to need for pre-deployed TH-2

Makes most sense for a base: surface infrastructure (habitat, mobility, power system) could be emplaced once and re-supplied

Marssurface

Marsorbit

TH-2TH-1SH

Transfer & Surface Hab and MOR

Earth-Mars transfer and Mars surface stay in TSH

Ascent to Mars orbit in MAV Return from Mars orbit to Earth in ERV This strategy could work both with and

without ISPP because the MAV crew compartment could be made very light (>5 mt)

Example: Mars semi-direct, Mars DRM 1.0, 3.0

The Mars staging orbits would likely be dependent on the utilization of ISPP:

ISPP: highly elliptic Mars staging orbit No ISPP: circular low Mars staging orbit

3 vehicle designs required: TSH, MAV, ERV

Mars surface and Mars orbit rendezvous

Split mission with pre-deployment is required in case of ISPP for MAV ascent propellants

Marssurface

Marsorbit

ERV MAV TSH

Mars Descent / Ascent Vehicle and MOR

Earth-Mars and Mars-Earth transfers in Interplanetary Transfer Hab (ITH)

Mars descent and ascent in Mars Descent & Ascent Vehicle (MDAV)

Option for propulsive abort to orbit exists during Mars powered descent

Surface stay in surface hab (SH) This strategy works both with ISPP and

without

The Mars staging orbits would likely be different for ISPP and no ISPP:


3 vehicle designs required: ITH, MDAV, SH Mars surface and Mars orbit rendezvous Makes most sense for a base: surface

infrastructure (habitat, mobility, power system) could be emplaced once and re-supplied

Marssurface

Marsorbit

ITHSH MDAV

Pre-Deployed Mars Descent / Ascent Vehicle and MOR


Mars descent in MDAV-2, Mars ascent in pre-deployed MDAV-1

Surface stay in pre-deployed surface hab (SH)

This strategy would only be chosen if ISPP is utilized for the MDAV

Mission mode requires either a single landing site or a chain of landing sites



3 vehicle designs required: ITH, MDAV, SH Mars surface and Mars orbit rendezvous Split mission required due to pre-deployed

MDAV Makes most sense for a base: surface

infrastructure (habitat, mobility, power system) could be emplaced once and re-supplied

Marssurface

Marsorbit

ITHSH MDAV-1MDAV-2

Landing & Surface Habitat and MOR


Mars landing & surface stay in Landing & Surface Hab (LSH)

Ascent to Mars orbit in MAV

This strategy works both with ISPP and without



3 vehicle designs required: ITH, LSH, ERV Mars surface and Mars orbit rendezvous

Marssurface

Marsorbit

ITHMAV LSH

Mapping of Major Mars Elements to Mission Phases

Major Mars Systems and Phases ES

->E

O

EO

EO

->E

MT

EM

T

EM

T->

MO

MO

MO

->M

S

MS

MS

->M

O

MO

MO

->M

ET

ME

T

ME

T->

ES

In-space and Surface Habitation x x x x x ? x ? x x xAscent Crew Compartment ? ?Crew Earth Ascent and Entry Vehicle x ? ? xEarth Launch and Departure Propulsion x xMars Aero-systems x xDescent and Landing Systems xMars Ascent and Departure Propulsion x xSurface Exploration Systems xSurface Power Systems xISRU Systems ?

Major Drivers of Selected Mars Systems

Earth Launch and Departure Systems are driven by the total payload which must be launched towards Mars and the Delta-V associated with the required Earth-Mars transfer trajectory.

In-Space Habitation Systems are driven by crew size and the duration of the Earth-Mars transfer trajectory

Mars Aero-Systems for Mars orbit capture and Mars entry and descent are driven by the payloads they must support and the Mars atmospheric entry velocities of the Earth-Mars transfer trajectories they must withstand

Mars Landing Systems are dependent upon the payload they must deliver to the surface and the state (velocity and altitude) at which they must begin operation

Earth Departure Delta-V Standard (no abort option) Conjunction Trajectories

3.50

3.60

3.70

3.80

3.90

4.00

4.10

4.20

4.30

4.40

4.50

4.60

4.70

4.80

4.90

5.00

120 130 140 150 160 170 180 190 200 210 220 230 240 250 260 270

Earth-Mars Transfer Time [Days]

Ear

th D

epar

ture

Del

ta-V

[km

/s] 2020

2022

2024

2026

2028

2031

2033

2035

2037

Earth Departure Delta-V Conjunction Trajectories with Earth-Mars-Earth Abort Option

3.5

3.6

3.7

3.8

3.9

4.0

4.1

4.2

4.3

4.4

4.5

4.6

4.7

4.8

4.9

5.0

0 0.3 0.6 0.9 1.2 1.5 1.8 2.1 2.4 2.7 3

Abort Delta-V [km/s]

Ear

th D

epar

ture

Del

ta-V

[km

/s]

2020

2022

2024

2026

2028

2031

2033

2035

2037

03-Year

Free Return2-Year

Free Return

Mars Entry Velocity

Abort Type

No Abort

Prop Abort or Hybrid

FRT Abort

Outbound: TSH

Outbound: ERV

Outbound: TSH or ERV

Outbound Abort? Ascent Abort? Outbound Config # Ares I / COTS # Ares V (Base) # EDS (Base) CEV (COTS)

Yes Yes MAV & ERV 0 3 (2) 3 (2) 1

Yes Yes TSH & ERV 0 3 3 1

Yes YesTSH, Earth

Aerocapture1 3 3 2 (1)

Yes NoTSH, Earth

Aerocapture0 3 3 1

Yes YesTSH &

Droppable CEV0 4 3 2

No Yes TSH 1 3 3 2 (1)

No No TSH 0 3 3 1

Abort Dependencies – MOR Architectures

Ares Launch Vehicle Capability

0

10,000

20,000

30,000

40,000

50,000

60,000

70,000

80,000

90,000

100,000

110,000

120,000

130,000

0 500 1,000 1,500 2,000 2,500 3,000 3,500 4,000 4,500 5,000

LEO Departure Delta-V [m/s]

To

tal

Pay

load

[kg

]

1 EDS, 0 mt added in LEO













1 EDS, Max added in LEO

2 EDS, HEO Rendezvous

2 EDS, Staged departure (max)

Single Launch

"1.5 Launch"1

2

35

4

Selected Mars Launch and Earth Departure Configurations

Ares VCore & SRBs

Ares V EDS

Blue = Payload; Gray = EDS Fairing; Red = Nuclear Thermal Stage

1 Ares V, H2/O2 2 Ares V, H2/O2 1 Ares V, NTR 2 Ares V, NTR

Launch and Earth Departure Performance

Trans-Mars Injection Performance vs. Transit Time (Minimum Across Mission Opportunities)

0

10

20

30

40

50

60

70

80

90

100

110

120

120 130 140 150 160 170 180 190 200 210 220 230 240 250 260 270

Earth Mars Transit Time [Days]

TM

I P

ay

loa

d [

mt]

2 Ares V, NTR

1 Ares V, NTR

2 Ares V, H2/O2

1 Ares V, H2/O2

ESAS Launch Vehicle Mars Capability 1 CaLV 1 CLV, 1 CaLV 2 CaLV, 1 EDS 2 CaLV, 2 EDS

~40 mt TMI ~50 mt TMI ~60 mt TMI ~90 mt TMI

~30 mt MO ~37 mt MO ~45 mt MO ~67 mt MO

~20 mt MS ~25 mt MS ~30 mt MS ~45 mt MS

TMI – Trans-Mars Injection; MO – Mars Orbit; MS – Mars Surface

Georgia Tech CE&R Aeroentry Analysis (Ventry = 4.63 km/s, L/D = 0.5)

15 m diameter aeroshell

Aeroshell Sizing Impact on TMI Mass CE&R aerocapture and aeroentry analysis indicated that aeroshell sizing

(diameter) would have a major impact on maximum mass of Mars systems ESAS CaLV has a fairing diameter of 8.4 meters, although larger fairings for Mars

systems would likely be possible For equal ballistic coefficient, the following entry mass limits likely apply for

entry systems of the specified diameter:

8.4^2 = 71 10^2 = 100 12^2 = 144 15^2 = 255

Diameter [m] 8.4 10 12 15

Entry Mass Range [mt] 25 to 31 35 to 44 51 to 64 80 to 100

10 m 15 m12 m8.4 m

Allowable Entry Mass vs. Aeroshell Diameter

0

20

40

60

80

100

120

8 9 10 11 12 13 14 15 16Aeroshell Diameter [m]

TM

I / E

ntr

y M

ass

[mt]

Mach 3 @ ~9 km Altitude

Mach 3 @ ~12 km Altitude

Aeroshell Sizing Impact on TMI Mass CE&R aerocapture and aeroentry analysis indicated that aeroshell sizing

(diameter) would have a major impact on maximum mass of Mars systems ESAS CaLV has a fairing diameter of 8.4 meters, although larger fairings for Mars

systems would likely be possible For equal ballistic coefficient, the following entry mass limits likely apply for

entry systems of the specified diameter:

8.4^2 = 71 10^2 = 100 12^2 = 144 15^2 = 255

10 m 15 m12 m8.4 m Diameter [m] 8.4 10 12 15

Entry Mass Range [mt] 25 to 31 35 to 44 51 to 64 80 to 100

1 LV, 1 EDS

1.5 LV, 1 EDS

2 LV, 1 EDS

2 LV, 2 EDS

Highly Elliptic Mars Orbits

Use of highly elliptic Mars orbits can decrease trans-Earth injection (TEI) delta-V relative to TEI from low Mars orbit

In this analysis, a ~10 hour period highly elliptic orbit is assumed in elliptic orbit cases, which decreases TEI delta-V by 1,000 m/s (this delta-V is added to Mars ascent requirements)

For the above chart, the pericenter is 3,700 km, equivalent to an altitude of ~300 km

Mars Elliptic Orbits (Delta-V from LMO and Period of Elliptic Orbit)

0

20

40

60

80

100

120

140

0 20,000 40,000 60,000 80,000 100,000 120,000

Apocenter [km]

Delt

a-V

[10 m

/s]

an

d P

eri

od

[h

]

Delta-V

Period

On Mars

Moon-Mars Thermal Comparison(Regolith Surface Temperature)

Moon-Mars-Boston Thermal Comparison(Regolith/Concrete Surface Temperature)

4500km4500km

Mars Water

Motivation for Solar Power on Mars

Mars missions studies frequently select nuclear fission reactors for providing surface power based upon technical considerations

e.g., mass, complexity, volume, etc. However, requiring nuclear power places a policy constraint on the

critical path for human Mars missions Need sustained funding to develop and test reactor Need political approval for every Mars mission including a reactor

Such a constraint compounds already existing policy challenges Increased dependence on the stance of political parties/politicians Sensitivity to changes in public view-point with respect to nuclear power

in general Other options such as dynamic isotope power systems also suffer

from approval and funding constraints

If solar power can be used as primary power source on Mars it could increase the political feasibility and sustainability of human Mars missions

Note: These policy constraints would also apply to nuclear thermal and nuclear electric propulsion

Daily Solar Incidence Energy Levels (Tracking Arrays, No Atmosphere)

0

2

4

6

8

10

12

0 100 200 300 400 500 600 700

Date in Sols (Perihelion = 0)

kW-h

(so

lar)

/ m

^2 /

sol

Equator

45-degrees North

45-degrees South

Energy Storage Factor Relative to Equator, Sized by Winter Solstice

0

0.5

1

1.5

2

0 5 10 15 20 25 30 35 40 45 50 55 60 65Latitude [deg]

Fac

tor

[-] Storage Capacity Factor

Is Solar Power Technically Feasible?

Preliminary analysis during our CE&R study indicated that solar power would be feasible as the primary power source for human Mars missions

Reasonably straightforward for missions without ISRU, more challenging for missions with extensive ISRU

We will conduct a further study of the technical feasibility and consequences of relying upon solar power for human Mars missions

Factors to be considered include required day and night-time power/energy levels, landing location (latitude), seasons and distance to Sun, atmospheric opacity and light scattering (including statistical nature of dust storms), array type (tracking, fixed-incline, fixed-horizontal)

Relying on solar power would likely place constraints on where missions could be conducted, available power levels, and may drive towards a single base (if a large investment is required to emplace the solar power system)

Winter Solstice Light and Dark Periods by Latitude

0

2

4

6

8

10

12

14

16

18

20

22

24

26

0 10 20 30 40 50 60 70 80 90

Latitude [deg]

Per

iod

[h

ou

rs]

Dark

Light

Energy Storage and Generation Factors Relative to Equator, Sized by Winter Solstice

0

1

2

3

4

5

6

7

8

9

10

0 5 10 15 20 25 30 35 40 45 50 55 60 65

Latitude [deg]

Fac

tor

[-]

Generating Capacity

Storage Capacity

Illumination Fraction over Martian Year for Selected Latitudes

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1

0 30 60 90 120 150 180 210 240 270 300 330 360

Solar Longitude, Relative to Winter Solstice [degrees]

Daily

Illu

min

atio

n Fr

actio

n [-]

0 Degrees

10 Degrees

20 Degrees

30 Degrees

40 Degrees

50 Degrees

60 Degrees

70 Degrees

80 Degrees

90 Degrees

Illumination Fraction over Martian Year for Selected Latitudes

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1

0 30 60 90 120 150 180 210 240 270 300 330 360

Solar Longitude, Relative to Winter Solstice [degrees]

Daily

Illu

min

atio

n Fr

actio

n [-]

33 Degrees

36 Degrees

39 Degrees

42 Degrees

45 Degrees

48 Degrees

51 Degrees

54 Degrees

57 Degrees

60 Degrees

Mars Water

Mars Elevation

Moon and Mars

What to do on Moon to prepare for Mars

Three main areas in which the Moon can offer aid in preparing for Mars, which could be considered objectives for lunar campaigns:

1. Testing systems, technologies, and procedures for Mars exploration in an environment distinct from Earth.

2. Increasing understanding of partial gravity (possibly coupled with radiation) impacts on crew health and performance.

3. Providing an intermediate milestone for human space exploration efforts.

How to measure those things

Metrics for each of the three areas of objectives for Mars preparation using the Moon are as follows:

1. Degree to which lunar systems are similar to Mars systems and fraction of Mars systems validated during lunar activities

2. Number of crew exposed to particular durations of lunar partial gravity (e.g., # at 1 month, # at 3 months, # at 6 months, etc.), with longer durations preferred

3. Date of initial occurrence of high visibility events (e.g., lunar vicinity flight, human lunar landing, long-duration mission, total surface time of Mars mission)

Mars Exploration Elements Following list of elements required for Mars exploration

Earth Launch and Entry Crew Cabin(s) Heavy Lift Launch Vehicle and Earth Departure Systems Descent Stage Heatshields Long-term Surface Habitat Mars Ascent Vehicle (Cabin and Propulsion) Earth Return Vehicle (Habitat and Propulsion) EVA and Mobility Systems Surface Power Systems

Following list of technologies beneficial for Mars missions In-Situ Propellant Production/In-Situ Consumables Production Mars ISRU Compatible Propulsion (e.g., CH4/O2, C2H4/O2)

Items denoted in blue indicate high potential for Moon-Mars commonality

Moon-Mars Exploration System Commonality

Shuttle Operations

Station Operations

Mars Exploration

System Development

Lunar Transportation System Operations

Lunar Long Duration System

Development

Budget Limit

Lunar Long Duration System Operations

Mars Exploration

System Operations

Lunar Transportation System Development

Shuttle Operations

Station Operations

Mars Exploration

System Development



Development

Budget Limit

Lunar Long Duration System Operations

Mars Exploration

System Operations


Shuttle Operations

Station Operations



Development

Budget Limit


Lunar Long Duration System Ops

Mars Exploration

System Operations

Mars Exploration

System Development

Shuttle Operations

Station Operations



Development

Budget Limit


Lunar Long Duration System Ops

Mars Exploration

System Operations

Mars Exploration

System Development

Shuttle Operations

Station Operations

Mars Unique Element

Dev.

Transportation System Operations

Long Duration System

Development

Budget Limit

Transportation System Development

Long Duration System Operations

Mars Unique Element Operations

Shuttle Operations

Station Operations

Mars Unique Element

Dev.

Transportation System Operations

Long Duration System

Development

Budget Limit

Transportation System Development

Long Duration System Operations

Mars Unique Element Operations

Common Moon-Mars Exploration System,Option to Maintain Lunar Missions

Distinct Moon, Mars Exploration Systems,Lunar Missions Curtailed

Distinct Moon, Mars Exploration Systems,Lunar Operations Maintained

If distinct systems are developed for Moon and Mars, we may:

Significantly delay Mars operations Need to curtail lunar operations to enable

Mars (development, operations), resulting in a Moon-Mars mission gap

Never get to Mars at all, because the renewed major investment is not sustainable

By developing a common Moon-Mars exploration system, we can overcome these obstacles and also:

Directly validate key Mars elements during lunar missions

Gain experience in routine production and system operation, decreasing cost and risk

Avoid workforce disruption during transition from Moon to Mars, and possibly continue lunar operations during Mars missions

Provide direct tie between Moon and Mars exploration in the eyes of the public and Congress

Commonality Strategy – Transportation Development Roadmap

Design Philosophy: Maximize hardware commonality to minimize gap between lunar and Mars missions and overall development and production costs

CEV + IPU (27 m3 ):

Integrated aeroshell

Mars Mission Hardware

LEO / ISS Mission Hardware

Common in-space propulsion stage (LCH4 / LOX):Core propulsion stageXL strap-on tanksXXL strap-on tanks (ERV)

Heavy Lift Launch Vehicle:(“2 stages”, 100 mt to LEO)

Short Lunar Mission Hardware

Habitat core and inflatablepressurized tent forplanetary surfaces:

Long Lunar Mission Hardware

Note: Block upgrades across phases are not depicted

LEO propulsion stage:

CEV launch vehicle:

CEV power pack:

LAT for CEV capsule:

SDLV upper stage (125 mt to LEO),potentially EDS-derived:

Mars landing gear & exoskeleton:

Engine 1 (LCH4 / LOX)Restartable, non-throttleable:

Common Earthdeparture stage(LH2 / LOX)

Engine 2 (LCH4 / LOX)Throttleable:

Lunar landing gear & exoskeleton:

Base Moon-Mars Exploration System Commonality Concept

High-level commonality concept developed during Base Period using selected Moon and Mars architectures

Commonality focused on design reuse of complete elements, with modularity in “Yellow Stage” and habitat design

Develop high-level scheme to identify elements where commonality may be beneficial Can be based upon elements with similar capabilities (or requirements) Need to be careful which requirements are compared

e.g., for a propulsion stage, the combination of delta-v, payload, and thrust characterize the capability (to first order); taken in isolation they do not

Develop commonality concept in further detail Trades must be performed between modularity/platforming or “stretchable” options relative to a single design for many use cases

Note: While commonality shown for a particular pair of architectures, approach is not unique to those chosen

Lunar Transportation Architecture Mars Transportation Architecture

Extensible Destination Vicinity Propulsion System

Core LCH4 / LOX propulsion stage

Derived LCH4 / LOX propulsion stages

Exploded viewIntegrated view

-4 non-throttleable LCH4/LOX engines-4 common-bulkhead tanks (CBH)-primary and secondary structure

Mars TEI Mars ascentLunar / Mars descent

-4 non-throttleable LCH4/LOX engines-4 common-bulkhead tanks-4 spherical extension tanks (2 CH4, 2 O2)-primary and secondary structure-dedicated exoskeleton for Mars TEI

-4 non-throttleable LCH4/LOX engines-8 common-bulkhead tanks-primary and secondary structure-additional primary structure

-4 throttleable LCH4/LOX engines-8 common-bulkhead tanks-primary and secondary structure-dedicated Moon and Mars landing gears and exoskeletons

Common Destination Vicinity Propulsion System

Modular solution for Destination Vicinity Propulsion System Common propulsion stage core employed in all use-cases (sized by Lunar Ascent & TEI) Duplicate set of tanks (relative to core) provides additional propellant for Lunar/Mars Descent

and Mars Ascent Extra-large set of strap-on tanks used for TEI from Mars on Earth Return Vehicle Descent stage structural ring and landing gear specific to destination due to distinct loading

conditions Common ascent engines, common descent engines for Moon [2 engines] and Mars [4 engines]

Moon MarsCrew Transport

Surface Habitat

Mars Ascent Vehicle

Transfer and Surface Habitat

Earth Return Vehicle

Moon-Mars Common System Vehicle Stacks

Post-Earth departure commonality mass overhead relative to customized systems:

Lunar Direct Return (Arch 1) Mars Orbit Rendezvous: Combined Trans. and Surf. Habs (Arch. 969)

Short Mission Long MissionLunar Crew

Transfer System

Lunar Long-Duration

Surface Habitat

Outbound Transfer & Surface Habitat

Earth Return Habitat & Propulsion

Mars Ascent Vehicle & Return CEV

81 100 112

112 106

106

106

106

59 39

36 34

9

21

9 mt

27 AS: 33

9

DS: 33DS: 33

Hab: 49

TEIS: 57

Hab: 25

HS: 34 HS: 34HS: 34

Number launches (HLLV+CEVLS):

Elements combine together to form vehicle stacks for variety of missions

Numbers at left represent wet mass in metric tonnes of elements in LEO

Earth Departure Stages have the same dry mass (11 mt) and maximum wet mass (112 mt)

CEVLV capacity 30 mt Lunar HLLV capacity 100 mt Mars HLLV upgraded to 125 mt

Low commonality overhead due to appropriate use of modularity to support variants

2+1 2+0 3+1 3+0 3+0

1% 2% 4% 3% 2%

IMLEO commonality overhead relative to customized systems: 13% 20% 4% 4% 3%

63% savings in unique element dry mass for common vs. custom system design

For modest mass increase, Mars-back commonality offers significant savings in development and production

Extending ESAS Elements to Mars Missions Based upon the significant capability of ESAS launch vehicles to deliver

payloads to Mars, options to extend the remaining elements were assessed

In the baseline architecture presented: Mars-dedicated aeroentry and propulsion systems are developed for

aerocapture/descent, landing, ascent, and trans-Earth injection The CEV is extended to provide habitation during Earth launch and entry, and during

Earth-Mars and Mars-Earth transit in combination with the LSAM-derived Mars Landing and Ascent Vehicle crew compartment, as part of a dual-launch, dual-heatshield crew transportation system

A large surface habitat capable of supporting up to 6 crew members is positioned to the Martian surface prior to the arrival of the crew in a single launch

Single launch logistics flights pre-position consumables, surface exploration equipment, and power infrastructure for initial mission, resupply consumables and spares for subsequent missions

Total number of CaLV launches required (no CLV launches are needed) is presented for a low launch demand and a high launch demand scenario

“Low demand” could be achieved with 4 crew (in one crew transportation system), methane-oxygen propulsion for maneuvers near Mars, and ISRU for both consumables production and propellant production

“High demand” could be achieved either with 4 crew (in two crew transportation systems), hypergolic propulsion, and no ISRU, or with 6 crew (in two crew transportation systems), methane-oxygen propulsion, and no ISRU

While not presented, intermediate launch demand options also exist with other crew size and technology combinations

Mars Crew Transportation System Concept

Launch Configuration Trans-Mars ConfigurationLogistics Flights Surface Habitat Crew Transport

Mars Orbit

Interplanetary Transfer

Earth Vicinity

Mars Surface

Conceptual Mars Exploration Architecture Based on Lunar Elements

Logistics Flights Surface Habitat Crew Transport

Conceptual Mars Exploration Architecture Based on Lunar Elements

Mars Orbit

Interplanetary Transfer

Earth Vicinity

Mars Surface

# of CaLVs by Mission

Initial Subs.Log. and Infra. 3 1Crew Transport 2 2

Total CaLVs 5 3

Initial Subs.Log. and Infra. 4 2Crew Transport 4 4

Total CaLVs 8 6

Low Demand

High Demand

Low demand: 4 crew, methane-oxygen prop, ISRU

High demand: 4 crew, hypergolic prop, no ISRU OR 6 crew, methane-oxygen prop, no ISRU

Mars Considerations for the Moona.k.a. Mars-Back to the Moon

Additional Considerations for building and extensible moon/Mars exploration system

What to do on Moon to prepare for Mars

Areas in which Moon can aid in preparing for Mars:

1. Testing systems, technologies, and procedures for Mars exploration in an environment distinct from Eartha) Gain operational and developmental experience;

test operations procedures

b) Test technologies for Mars systems (e.g., method of water recycling

c) Test Mars subsystems (e.g., ECLSS, fuel cell)

d) Test Mars systems (e.g., habitat, rover)

2. Increasing understanding of partial gravity (possibly coupled with radiation) impacts on crew health and performance

3. Providing an intermediate milestone for human space exploration efforts

Metrics for each area:1. Degree to which lunar systems,

subsystems, etc. are similar to Mars systems and fraction of Mars systems validated during lunar activities (with validated earlier being better)

2. Number of crew exposed to particular durations of lunar partial gravity (e.g., # at 1 month, # at 3 months, # at 6 months, etc.), with longer durations preferred

3. Date of initial occurrence of high visibility events (e.g., lunar vicinity flight, human lunar landing, long-duration mission, total surface time of Mars mission)

Measuring Readiness for Mars

It may be useful to define a “Mars Exploration Readiness Level” (MERL) to determine Mars preparedness

Would serve a similar function to the TRL scale for measuring technologies

Having such a scale would enable: Reviewing status Monitoring progress Determining criteria for tests

Lunar testing would be one means of increasing MERL Laboratories, terrestrial Mars-analogs, LEO, deep-space,

and Mars could also serve as test environments

MERL could be applied to areas 1 and 2 from prior listing (i.e., technical readiness and human physiological readiness)

MERL could be a hierarchical measure, and at top level would depend upon overall Mars architecture(s) under consideration

ME

RL

Time

Level Required by NASAto commence Mars missions

Potential Areas of Moon-Mars Commonality Habitat sub-systems Habitat structure

Note: Mars mission will require relatively large habitat volume, should limit surface assembly due to safety considerations

Mars EDL system would place additional constraints on habitat dimensions, CG Surface exploration systems (i.e., rovers, EVA aids, etc.) EVA suits

Possible differences in thermal, note also different in-situ resources may impact design Power systems (generation and storage)

For solar, storage is easier on Mars, generation harder Operations commonality (time lag and associated impacts)

Should lunar missions operate in this manner from the start? Resupply and logistics? Propulsion?

Ascent, TEI: LCH4/LOX, C2H4/LOX, RP-1/LOX?, etc. H2/LOX for descent?

Note: Lunar ISRU is intentionally not on this list

While these areas offer potential for commonality, the list does not indicate what should be done in designing lunar systems to make them common with Mars

Mars Considerations for Lunar Systems

In order to determine Mars-related considerations for the development of lunar systems (so as to aid in lunar system design decision making), we ask:

How is Mars different from the Moon?

How do the differences tend to impact Mars exploration systems?

What can be done with lunar systems to have higher relevance towards Mars systems?

Differences between Mars and the Moon (1)

Mars has an atmosphere Mars systems require an aeroshell for EDL

Aeroshell must be launchable at Earth (constrains diameter) Integrated system must be stable through atmospheric flight Tends to lead to compact, low cg descent stage and payloads for Mars For Moon: Consider designing payloads to meet Mars aeroshell constraints

CO2 is easily available on Mars Mars systems would tend to use atmospheric-based ISRU, rather than more

challenging, complex, and risk regolith-based ISRU For Moon: Consider use of Mars ISRU related processes in ECLS and power

systems (e.g., water electrolysis, Sabatier) For Moon: Do not justify lunar regolith or water-ice based ISRU on the basis

of relevance to Mars; do not predicate Mars ISRU on success with lunar ISRU

Mars surface radiation environment more benign than Moon Mars systems may require less shielding while on surface For Moon: Unclear if this carries implications; may impact common mobility

system design in that Mars systems would not need as much shielding

Differences between Mars and the Moon (2)

Campaign- / mission-level differences Mars missions will be longer and possibly feature larger crew

Required surface habitat volume will be greater for Mars than Moon For Moon: Consider oversized (relative to lunar requirements) habitat or

means for increasing habitat volume between Moon and Mars

Full Mars habitation capabilities are required from the start Mars systems would tend to limit assembly required to provide crew

habitation (assembly might be used to augment capabilities) For Moon: Consider means of providing core crew habitation functionality

without assembly

Mars missions do not have any opportunity for re-supply Leads towards pre-positioning of required consumables and sufficient up-

front planning to anticipate full needs of mission For Moon: Consider logistics strategy that does not rely on rapid re-supply,

but uses pre-placement or pre-planned re-supply

Differences between Mars and the Moon (3) Mars has approximately twice the gravity of the Moon

Mars surface operations will be under a higher gravitational load than lunar surface operations Tends towards desire for lighter space suits for Mars Tends towards decreased reliance on astronauts lifting, carrying items Mobility energy requirements will be higher for a given mass Mobility systems may become stuck more easily / more stuck on Mars For Moon: Consider means to decrease space suit mass further than may be needed for

lunar ops; limit degree to which astronauts lifting is required For Moon: Aim to minimize mobility “payload” mass; consider means to recover higher

weight mobility systems

Mars is further away from Earth than the Moon Mars missions will feature a significant round-trip communications lag

Mars missions operations will conducted in very different manner (relative to operations to date) in order to accommodate time lag; crew will have significantly greater autonomy, with Earth supporting and advising on a longer time-scale

Mars systems will require greater automation in order to monitor system health and limit workload on surface crew

For Moon: Consider means of simulating Earth-Mars communications lag in lunar operations; establish lunar surface operations methodology including time lag

Differences between Mars and the Moon (4) Mars is further from Sun and has a ~24 hour day

Solar incidence is lower although eclipse time is shorter Solar power generation more challenging on Mars Energy storage (e.g., regen fuel cells) is easier on Mars For Moon: Consider means to modularize lunar solar power generation system so that

additional units can be included in Mars case; consider means to decrease energy storage capacity for Mars relative to Moon

Note: Nuclear power is not necessarily required for Mars

Mars thermal environment is different from the Moon Regolith temperature more moderate on Mars than Moon, also have

atmospheric interactions on Mars Different thermal systems might be required between Moon and Mars (although more

analysis needed) For Moon: Identify how relevant lunar systems are towards Mars, determine options for

transitioning from lunar thermal environment to Mars

Mars may have or may have had indigenous life Should restrict contamination of sites of scientific interest with

organisms transported by humans; limit exposure of crew to martian organisms Techniques for sterile sampling of biologically interesting sites may be required Techniques for preventing biological transfer into crewed areas may be required For Moon: Consider testing of such techniques as part of lunar operations

Concept to implement delay in Earth-Moon communications so that lunar surface missions are conducted in a manner and with an organization that is relevant to Mars

Delay would be in Earth to Moon comm. leg (not Moon to Earth), so that mission support could actively assess situation and switch to real-time comm. in an emergency

Primary role of mission support organization would be to assist crew and monitor longer time-scale processes – real time decision-making would reside on the lunar surface

Option to include an Earth based “artificial artificial intelligence” organization may be worthwhile to reduce need to develop onboard automation prior to lunar missions

Would be staffed by personnel simulating the role of “automation”

Should be tightly tied to organization developing automation software, to inform the development of software and test various techniques

Human involvement in this area would be decreased / eliminated over time

Could also have 2 “virtual astronauts” on Earth with real-time comm. links

Could assist in difference in crew size between Moon and Mars

Could monitor EVAs, teleoperate rovers, etc. Could test various roles for crew members in a Mars

mission

Mars-Driven Lunar Missions Operations Concept

MissionSupport

4 to 20 min. lag

4 to 20 min. lag

Concept to implement delay in Earth-Moon communications so that lunar surface missions are conducted in a manner and with an organization that is relevant to Mars

Delay would be in Earth to Moon comm. leg (not Moon to Earth), so that mission support could actively assess situation and switch to real-time comm. in an emergency

Primary role of mission support organization would be to assist crew and monitor longer time-scale processes – real time decision-making would reside on the lunar surface

Option to include an Earth based “artificial artificial intelligence” organization may be worthwhile to reduce need to develop onboard automation prior to lunar missions

Would be staffed by personnel simulating the role of “automation”

Should be tightly tied to organization developing automation software, to inform the development of software and test various techniques

Human involvement in this area would be decreased / eliminated over time

Could also have 2 “virtual astronauts” on Earth with real-time comm. links

Could assist in difference in crew size between Moon and Mars

Could monitor EVAs, teleoperate rovers, etc. Could test various roles for crew members in a Mars

mission

Mars-Driven Lunar Missions Operations Concept

MissionSupport

Artificial Automation

Virtual Astronauts

8 to 40 min. lag

emer

gency

com

m.

human mars missions

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