lunar educational wide imaging satellite

University of Southampton

SESA3024 and SESA6055 - Mission to the Moon

LEWIS - Lunar Educational WideImaging Satellite

Alexander Godfrey

Lukas Gossnitzer

James MacCalman

Ahmed El Maghraby

Alessandro Melis

Nicole Melzack

Reetam Singh

Nikolay Tenev

Aurelien Toussaint

Maria Zaretskaya

April 24, 2013

2

Abstract

The Lunar Educational Wide Imaging Satellite (LEWIS) is designed with a

primary mission objective to image the lunar surface for education and outreach

purposes. Secondary objectives are the detection of atmospheric dust, the

analysis of the lunar radiation environment and the characterisation of the

structure of the regolith. The mission lifetime in the lunar orbit is required to

be at least six months and the launch mass must not exceed 300 kg. After being

placed in a geostationary transfer orbit (GTO) by the Ariane 5 launcher, three

different lunar transfer methods are contrasted: a direct chemical transfer, a

chemical transfer via a weak stability boundary point and an electric propulsion

low thrust transfer. A preliminary spacecraft design study is conducted for each

transfer option using a concurrent design environment. The level to which the

designs fulfill the mission objectives, or exceeds them, is evaluated. The final

recommendation for the orbiter is to use the weak stability boundary method.

The main factors leading to this decision include the relatively short transfer

time (80 days) which will minimise any degradation and debris impacts to the

spacecraft, the extended mission lifetime of 1350 days (compared with the 298

days offered by the direct chemical method) and the final orbit plane which is

only 4◦off the ideal 0◦argument of periselene.

CONTENTS 3

Contents

1 Introduction 6

1.1 Mission Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6

1.2 Science Objectives . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6

1.3 Spacecraft System Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7

1.4 Overview of Possible Transfer Methods . . . . . . . . . . . . . . . . . . . . . . . . . . 7

1.4.1 Direct Chemical . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7

1.4.2 Weak Stability Boundary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8

1.4.3 Low Thrust . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8

2 Payload Specifications 9

2.1 Primary Payload . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9

2.2 Secondary Payload . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10

2.2.1 Radiation Assessment Detector (RAD) . . . . . . . . . . . . . . . . . . . . . . 10

2.2.2 Spectrometer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11

2.2.3 Dust Detector . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12

3 Concurrent Design Approach 12

4 Direct Chemical Transfer Method 15

4.1 Mission Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15

4.1.1 Transfer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15

4.1.2 Final Orbit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15

4.2 Chemical Propulsion Subsystem . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17

4.2.1 Propulsion Subsystem Sizing . . . . . . . . . . . . . . . . . . . . . . . . . . . 17

4.2.2 Propulsion Subsystem Configuration . . . . . . . . . . . . . . . . . . . . . . . 18

4.3 Power Subsystem . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22

4.3.1 Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22

4.3.2 PCDU . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22

4.3.3 Battery Sizing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 23

4.3.4 Solar Array Sizing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 24

4.3.5 Solar Array Configuration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 25

4.3.6 Power Profile . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26

4.4 On-Board Data Handling Subsystem (OBDH) . . . . . . . . . . . . . . . . . . . . . 26

4.4.1 Overview . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26

4.4.2 Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 27

4.4.3 Components . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 27

4.4.4 Software . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 31

4.4.5 Considerations for Direct Chemical . . . . . . . . . . . . . . . . . . . . . . . . 32

4.5 Communication Subsystem . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 33

4.5.1 Link Budget . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 33

4.5.2 Link Quality . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 33

4.5.3 Ground Stations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 34

4.5.4 Carrier Frequencies . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 34

4.5.5 Losses . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 35

4.5.6 Component Selection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 35

4.5.7 Investigating Communications Windows . . . . . . . . . . . . . . . . . . . . . 38

4.5.8 Contingencies . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 39

4.5.9 Final Results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 40

4 CONTENTS

4.6 Attitude Determination and Control Subsystem (ADCS) . . . . . . . . . . . . . . . . 40

4.6.1 Control Modes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 40

4.6.2 Selection of Attitude Control Method . . . . . . . . . . . . . . . . . . . . . . 41

4.6.3 Quantifying the Disturbance Environment . . . . . . . . . . . . . . . . . . . 42

4.6.4 Selection and Sizing of ADCS Hardware . . . . . . . . . . . . . . . . . . . . . 43

4.6.5 Hardware Application and Resultant Thruster Requirements . . . . . . . . . 45

4.7 Thermal Control Subsystem . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 46

4.7.1 Primary Assumptions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 46

4.7.2 Method of Calculations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 48

4.7.3 Equilibrium and Eclipse Temperatures . . . . . . . . . . . . . . . . . . . . . 50

4.7.4 Thermal Control Decisions . . . . . . . . . . . . . . . . . . . . . . . . . . . . 51

4.8 Structure and Configuration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 52

4.9 Discussion of Budget Evolution and Sub-system Trade-offs . . . . . . . . . . . . . . . 55

4.10 Final Mission Budget . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 57

5 Weak Stability Boundary Transfer Method 58


5.1.1 Transfer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 58

5.1.2 Final Orbit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 58

5.2 Chemical Propulsion Subsystem . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 59

5.3 Power Subsystem . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 61

5.3.1 Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 61

5.3.2 PCDU . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 62

5.3.3 Battery Sizing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 62



5.3.6 Power Profile . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 62

5.4 OBDH . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 62


5.6 ADCS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 63

5.6.1 Control Modes and Selection of Attitude Control Modes . . . . . . . . . . . . 63

5.6.2 Quantifying the Disturbance Environment . . . . . . . . . . . . . . . . . . . . 63

5.6.3 Selection and Sizing of ADCS Hardware . . . . . . . . . . . . . . . . . . . . . 63

5.6.4 Hardware Application and Resultant Thruster Requirements . . . . . . . . . 64

5.7 Thermal Control Subsystem . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 65






5.9 Discussion of the Budget Evolution due to Sub-system Trade-offs . . . . . . . . . . . 68


6 Low Thrust Transfer Method 71


6.1.1 Transfer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 71

6.1.2 Final Orbit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 72

6.2 Electric Propulsion Subsystem . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 73

6.2.1 Low Thrust Reaction Control System . . . . . . . . . . . . . . . . . . . . . . 76

CONTENTS 5

6.3 Power Subsystem . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 76

6.3.1 Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 76

6.3.2 PCDU . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 77

6.3.3 Battery Sizing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 77



6.3.6 Power Profile . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 78

6.4 OBDH . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 78


6.6 ADCS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 79

6.6.1 Control Modes and Selection of Attitude Control Mode . . . . . . . . . . . . 79

6.6.2 Quantify the Disturbance Environment . . . . . . . . . . . . . . . . . . . . . 79

6.6.3 Selection and Sizing of ADCS Hardware . . . . . . . . . . . . . . . . . . . . 80

6.6.4 Hardware Application and Resultant Hardware Requirements . . . . . . . . . 80

6.7 Thermal Control Subsystem . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 81






6.9 Discussion of the Budget Evolution due to Sub-system Trade-offs . . . . . . . . . . . 86


7 Evaluation and Comparison of Transfer Methods 90

7.1 Transfer time . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 90

7.2 Final orbit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 90

7.3 Extended Mission Lifetime . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 90

7.4 Mass . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 91

7.5 Risk Assessment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 92

8 Observation Strategy 93

9 Final Choice of Orbiter 96

References 97

6 1 INTRODUCTION

1 Introduction

”I think we’re going to the moon because it’s in the nature of the human being to face

challenges. It’s by the nature of his deep inner soul... We’re required to do these

things just as salmon swim upstream.” - Neil Armstrong

The Lunar Educational Wide-angled Imaging Satellite (LEWIS) meets the new challenge

of acquiring high-resolution images of lunar surface for public relation and education

outreach purposes. The secondary instruments on the space-craft would perform

new scientific measurements which would deepen our knowledge of lunar science and

contribute towards the future human exploration of the moon whilst having a better

understanding of the lunar environment. LEWIS is scheduled to be launched in 2013

into a GTO onboard an Ariane 5 launch vehicle.

1.1 Mission Requirements

• Place the spacecraft in an appropriate lunar orbit

• Acquire images of the Moon and transmit them back to Earth

• Perform scientific measurements relevant to objectives outlined below

1.2 Science Objectives

The primary objective for LEWIS is to image lunar surface with a high resolution

camera (Remote sensing) in a visible spectrum region for education and outreach

purposes. It is desirable to collect a variety of visually different pictures of the moon.

The satellite will therefore collect images of the lunar surface from a variety of points

of view, including the earth whenever possible for its aesthetic property. The visual

effects of the four lunar eclipses occurring in 2014 and 2015 on the lunar surface will

also be observed. When possible, the view of the earth eclipsing the sun will also be

captured. The secondary science objectives are specified below:

• Investigate and characterise the radiation environment on lunar surface as well

as in the lunar orbit. This would be helpful in determining the biological effects,

caused due to exposure to radiation in the lunar environment for future lunar

manned missions as well as undertake a feasibility study of moon as a base for

deep solar missions.

• Undertake chemical and mineralogical mapping of lunar surface for distribution

of elements within the bandwidth of 1.0- 10.0 KeV (elements like Aluminium,

Magnesium, Silicon, Calcium, Iron and Titanium). Also perform mapping of

heavy radioactive elements like Radon, Uranium and Thorium with a high

spatial resolution. This would facilitate the study of moon as a base for mining

of scarce radioactive elements for future missions.

1.3 Spacecraft System Requirements 7

• Mapping of water on lunar poles. According to the research of Arnold[1979],

Iron reduces in the presence of solar wind protons to liberate water which ac-

cumulated at the polar cold traps in the permanent shadow in the form of ice.

Assuming that it is not indigenous and the implanting solar wind protons are the

only reducing agent for , we can investigate the presence of water on lunar caps

through the detection of Fe 3+ molecules. Hence Polar regions are considered

for the purpose of water mapping, as there is an absence of solar wind flux

in these areas due to which it has higher probability of the presence of water

molecules.

• Characterise the dust environment around the lunar orbit. This would allow

us to simulate better environment for lunar entry/re-entry of future mission

modules.

• Characterise the lunar plume and ejecta composition from previous impactor

missions to moon and undertake a study for lunar surface composition. Plume

from missions around lunar poles would facilitate this goal.

1.3 Spacecraft System Requirements

• Achieve variety in points of view by having an eccentric orbit. Abiding with

guidelines given by the customer, polar orbit with periselene of 100 km and

aposelene of 3600 km is chosen. To have the earth visible with the moon at

variable altitudes, argument of periselene of 0◦is desirable.[2]

• Total spacecraft mass not to exceed 300 kg and to fit within a volume constraints

within a 1666 mm height 1640 mm diameter cylinder. This is for the spacecraft

to fit in a launch vehicle specified by the customer. [2]

• Should be able to structurally resist the vibrations from the Ariane 5 launch

environment. [2]

1.4 Overview of Possible Transfer Methods

1.4.1 Direct Chemical

The direct chemical option accomplishes the transfer by executing two impulsive

burns: one to raise the apogee to the lunar orbital radius, and one for capture.

A small burn is performed amid course to achieve desired inclination and periselene

altitude.

The transfer fuel-mass requirement for this transfer is the highest amongst the

options discussed in this paper, scoring 100.1kg. This, however, is not far behind the

second highest, which is 96.6kg for WSB Option. In turn, the transfer time is found

8 1 INTRODUCTION

to be the shortest, from four to five days.

Calculations for this option has been done by two independent patched conic analyses

and an STK simulation. Ultimately, the values from STK has been used to design

the spacecraft.

The overall risk facing this option is the lowest amongst the options, as shown in the

end of this report.

1.4.2 Weak Stability Boundary

This type of transfer gains some of its ∆V requirement from the sun’s gravitational

perturbence by sending the orbiter to one of the Sun-Earth Lagrange points. For this

reason, the requirement is slightly less than that for direct chemical option. In turn,

the transfer time is more than ten-folds, scoring 80 to 100 days, as shown later in this

paper.

Two types of this transfer are considered: the conventional WSB transfer, in which

chemical propulsion is used to raise the apogee directly to L1, and a WSB transfer

via a lunar flyby, in which the apogee is only raised to the lunar orbit, whereby lunar

gravitational assist takes the orbiter to L1. For reasons outlined later in this report,

one with lunar flyby has been chosen for the mission.

Calculations for this option is done solely on STK.

1.4.3 Low Thrust

This option employes an electrical propulsion system with high specific impulses,

ranging from 500 to 5000 seconds. Due to the low thrust in the order of milliNewtons,

impulsive burn approximations are not applicable for fuel-mass estimations. For this

reason, the EMT software written by Dr. Hugh Lewis has been used to estimate the

∆V requirement and the transfer time. As shown in later sections the transfer time

for this option ranges from 200 days to two years.

The electric engines require power from 0.5 to two kiloWatts, which are higher

than the rest of the spacecraft subsystems’ needs combined. To accommodate these,

larger deployed solar panels are required.

This has allowed, however, the use higher-power RF transmitter to retrieve 1.7

times more data per orbit compared to the other options.

The risks associated with this option is the highest amongst the options, as shown in

the end of this report.

9

2 Payload Specifications

The payload instruments have evolved over the course of the projects further research

produced more suitable instruments for the mission. RAL space has given advice on

the primary instrument selection which provided a range of potential cameras suitable

for the mission. The secondary instruments have also changed as better alternatives

were found.

2.1 Primary Payload

The primary payload consists of two cameras on board the spacecraft. One camera

has a narrow field of view (FOV) of 3◦and the second camera has a wide field of view

(50◦). The reason for this choice is to have a range of images from the Moon and its

environment. To capture the interest of the public, images which show the Moon as a

whole and detailed close up images of its surface should be produced. Another reason

a wide field of view camera was chosen was to capture a full image of the Moon during

the four total lunar eclipses occurring in 2014 and 2015.

For the narrow FOV camera it was decided to have a camera with a high resolution

to capture detail. A number of cameras were compared based on this requirement

and it was found that the camera designed for the European Student Moon Orbiter

(ESMO) is the best solution. The ESMO camera is designed for a field of view of

3 and has more pixels than the other cameras which were considered; which makes

it the most suitable camera for capturing high resolution images. The images are

large in size (100 MB) in comparison to other cameras (e.g. BADR-B camera). This

will have to be taken into account by communications as 10 images per orbit will be

taken. The ESMO camera is also the best choice as it is lighter and has a lower power

consumption in comparison to the other cameras considered.

For the wide FOV camera the resolution requirement is not as high. The camera

must produce 10 pictures a day. As these pictures will be of smaller size (44 MB),

they will not put strain on the communications. The BADR-B CCD camera is a wide

angle camera, launched in 2001 on board the Earth observing BADR-B satellite to

detect the presence of small clouds to aid with the rejection of cloud contaminated

data.[3] The current model has a field of view of 8.5◦but it was assumed that the

optics inside the camera can be changed to allow a field of view of 50◦. The camera

has around half the pixels of the ESMO camera which translates into a smaller image

size (44MB). The relatively smaller image size will allow the communications to cope

with the heavy load of data. The specifications for each camera is shown in Table

2.1.1.

10 2 PAYLOAD SPECIFICATIONS

Camera ESMO BADR-BMass 1.93 kg 2.5 kgPower (Operating) 2.8W 10 WImage Size 100MB 44MBFOV 3◦ 50◦

Exposure Time 87ms 20ms

Table 2.1.1: Specifications for the ESMO and BADR-B Cameras

2.2 Secondary Payload

2.2.1 Radiation Assessment Detector (RAD)

The Radiation Assessment Detector (RAD) is an energetic particle detector designed

to measure a broad spectrum of energetic particle radiation. [4] It is a lightweight

and energy efficient passive detector which acquires radiation data from Galactic

Cosmic Rays (GCRs) and ionised particles from Coronal Mass Ejection (CMEs). The

acquired information will be used to assess the potential radiation hazard for future

Moon man-missions and Moon based colonies, and how that radiation dosage affects

the spacecraft subsystems during the transfer and the six month Moon mission.

The RAD combines both charged and neutral particle detection capability over a

wide dynamic range in a compact, low mass, low power instrument.[4] These capa-

bilities are required in order to measure all the important components of the radiation

environment.[4]

(a) Schematic of the RAD Sensor [4] (b) Cross section of RAD [5]

Figure 2.2.1: Schematics of Radiation Assessment Detector

2.2 Secondary Payload 11

The RAD consists of two main parts: the RAD Sensor Head (RSH) and the RAD

Electronic Box (REB) integrated in one container.[4, 5, 6] Specifications of the RAD

are shown in Table 2.2.1, and the RAD’s schematics are in Figures 2.2.1a and 2.2.1b.

Parameter ValueMass 1.56 kgPower (Operating) 4.2 WData Volume 400 kB/day

Table 2.2.1: Specifications for the Radiation Assessment Detector(RAD)

2.2.2 Spectrometer

The Lunar Exploration Analysis Group (LEAG) have set ”Characterizing the structure

and layering of the regolith” as one of their objectives. [7] This can be achieved

by selecting an appropriate spectrometer. The Chandrayaan-1 X-ray spectrometer

(C1XS) is designed to measure absolute and relative abundances of major rock-

forming elements (principally Mg, Al,Si, Ca, Ti and Fe) in the lunar crust with

spatial resolution 25 km. [8] The C1XS spectrometer was designed by the Rutherford

Appleton Laboratory (RAL) for the Indian Space Research Organisation (ISRO)

Chandrayaan-1 lunar mission and launched in 2008. [8]

The C1XS spectrometer has a higher power consumption and mass than other

spectrometer we looked at: the Mars Odyssey Neutron Spectrometer and the Mercury

Neutron Spectrometer. However, it is capable of performing measurements on a wider

range of elements. The CAD image of the C1XS spectrometer is shown below.

Figure 2.2.2: CAD image of the C1XS Instrument showing coalligned front detectors,deployable radiation shield and 140◦field-of-view.

12 3 CONCURRENT DESIGN APPROACH

Parameter ValueMass 5.56 kgPower (Operating) 25.5 WData Volume 36 MB/orbitSpatial Resolution 25 km

Table 2.2.2: Specifications for the The Chandrayaan-1 X-ray Spectrometer (C1XS)

2.2.3 Dust Detector

The dust environment of the Moon and its fragile atmosphere are of great interest to

the scientific community. The Lunar Exploration Analysis Group (LEAG) describes

this area of study as a key research theme for future missions. [7] A Piezo Dust

Detector (PDD) will perform consistent dust monitoring to better understand dust

migration patterns on the Moon by direct detection of particle impacts. The PPD is

a modular, miniaturised in-situ measurement device. The modular design allows an

addition of detector units to increase the sensor surface or measure impacts on multiple

spacecraft surfaces. [9] The detector has a low mass, low power consumption, low data

rate and small size. This flexible design makes the PDD easy to accommodate on the

spacecraft.

The detector will provide physical parameters of impacting dust and debris particles

such as velocity, mass and impact energy. [9] The size of detectable particles will be

in the range of 1 µm to 1 mm at a velocity of up to 10 km/s. [9]

Parameter ValueMass 0.5 kgPower (Operating) 3 WData Volume 36 MB/orbit

Table 2.2.3: Specifications for the Piezo Dust Detector (PDD)

3 Concurrent Design Approach

Every effort was made by the team to work concurrently throughout the design

process. All decisions were thoroughly discussed and made as a team. The team

was faced with a decision on whether to use the ESA SCDE concurrent design system

or GoogleDocs. It was decided that GoogleDocs is more adequate for the reasons

explained below.

The design team is 10 people. This is a large and international group and in order

to work concurrently, the team has to be flexible to work anywhere and anytime.

13

Using the GoogleDocs allows remote access to the spreadsheets so even members who

are working from remote locations can edit the spreadsheets and contribute to the

design process.

Using GoogleDocs makes the design process more dynamic as the changes to the

configuration of subsystems are immediately visible to the each subsystem engineer.

This increases the number of iterations and optimizes the design faster. This also

makes it easier to experiment with subsystems as the system level implications are

immediately visible.

Errors can be resolved a lot faster by using GoogleDocs rather than the ESA

SCDE system. Everyone can see the spreadsheets in GoogleDocs in real time and all

subsystems are always connected. This allows mistakes to be spotted immediately in

contrast to the SCDE system when the subsystems are merged every few days.

Figure 3.0.3: Budgets tab in the concurrent design spreadsheet

One of the problems with GoogleDocs was that is was easy to fall into circular

references. To resolve this problem, one segment of the circular chain was cut and

manually equated while manually keeping track of the error in the approximation.

Another way to deal with this was by automating the process by writing scripts

which size the system iteratively. This was done for the propulsion subsystem sizing

14 3 CONCURRENT DESIGN APPROACH

for the direct chemical and weak stability boundary methods.

In Figure 3.0.3 is an example of a spreadsheet. The user can see when the last edit

was made and by whom. The current users are displayed in the top right hand corner.

The tabs at the bottom show other spreadsheets. Spreadsheets are interconnected by

linking values.

Figure 3.0.4: Attitude control subsystem spreadsheet construction

Figure 3.0.5: The implemented designprocess loop.

Figure 3.0.4 shows how each sub

system spreadsheet was structured. The

Inputs section is linked to other engineers’

Outputs section. Below the Inputs

section is a Calculation section where

all the necessary computations are done.

On the right of the spreadsheet is a

Requests section where requests can be

made by other engineers. The Outputs

section displays all the values needed by

other engineers.

To make continuous optimisation iterations as a group we used the design process

shown in Figure 3.0.5. All decisions were discussed as a group. After the discussion

changes were implemented. Then a system analysis of the design was done. Cor-

rections were made as problems arose. Then the group discussed the new iteration.

15

4 Direct Chemical Transfer Method

4.1 Mission Analysis

4.1.1 Transfer

To determine the transfer ∆V requirement for LEWIS, patched conics method was

used as a first estimate in Microsoft Excel, results shown in Table 4.1.1. It was shown

that transfer ∆V is minimized when trans-lunar ejection is performed at GTO Perigee

and lunar encounter occurs when the moon is at its Apogee. Another independent

patched conics approximation was performed in MATLAB, producing transfer ∆V of

1155 m/s. To obtain a firm result, STK analysis was performed. It was found that in

December 2013, the intersection between lunar orbit and the Earth’s equatorial plane,

which is close to LEWIS’ initial orbital plane, occurred at its highest position in the

year 2013. The transfer consists of three burns: TLE, a mid-course correction burn

is then performed to achieve the desired inclination and periselene altitude (PA) of

90◦and 100 km, respectively. LEWIS is then allowed to coast to the periselene, where

TLI is performed to achieve the desired aposelene altitude (AA). The resultant ∆V

requirement has been found to be 1194 m/s, with 5 day transfer time.

ConfigurationTLE ∆V TLI ∆V Total Transfer ∆V Transfer Time

[m/s] [m/s] [m/s] [days]Perigee to Perigee 693.7 498.9 1192.6 4.58Perigee to Apogee 704.4 466.5 1170.9 5.40Apogee to Perigee 2457.9 412.7 2870.6 5.24Apogee to Apogee 2480.7 396.3 2877.0 6.10

Table 4.1.1: Transfer∆V Results using Patched Conics Approximation

4.1.2 Final Orbit

1

2

3

4

30

210

60

240

90

270

120

300

150

330

180 0

Orientation of Final Orbit for Direct Chemical Option

Orbital Radius [Moon Radius]

South - North

LEWIS Orbit

Figure 4.1.1: Final Orbit Achieved bydirect chemical Option. AoP = 320.6◦

This choice of transfer results in a

final orbit with the highest Argument

of periselene (AoP) amongst the three

options described in this report, as shown

in Figure 4.1.1. Although it is seen from

the spacecraft system requirements that

AoP of 0◦is desirable, it was decided that

no further maneuvres to obtain AoP of

zero will be performed, since similar pho-

tographs can be taken from this orbit,

and further maneuvres would increase

the ∆V requirement by upto a factor of

1.5.

16 4 DIRECT CHEMICAL TRANSFER METHOD

The near-45◦AoP results in strong gravitational orbital perturbation, resulting in

the highest stationkeeping requirement amongst the three options, of 340m/s/year.

Stationkeeping To achieve minimum mission duration of 6 months, the PA must

be maintained well above surface. The station-keeping strategy employed for this

option is to boost the PA and AA when either PA or AA is altered by tolerance of

10 km. The inclination, Ascending Node and AoP of the orbit are not controlled, as

controlling these would require large ∆V , and maintaining these offers no benefit to

the Science Objectives, as long as orbiter stays above the surface with high variation

in orbital altitude.

Simulation in STK shows that this strategy requires 340m/s of ∆V over 1 year

since Translunar Injection (TLI); the resultant Altitude over 1 year since TLI is shown

in Figure 4.1.2. Using all available propellant on board, the designed orbit can be

maintained for 9 months since TLI. Using measured data of the lunar radiation en-

vironment, the solar array was sized to provide minimum power at this time, as

discussed in Section 4.3.

Figure 4.1.2: STK simulation on PA and AoP of LEWIS when the orbit is activelycontrolled. Keeping PA at 100 km ± 10 km for 1year requires 340m/s. By linearinterpolation given the fuel carried, LEWIS will keep the station for 9 months inlunar orbit.

4.2 Chemical Propulsion Subsystem 17

4.2 Chemical Propulsion Subsystem

The chemical propulsion subsystem generates the thrust for orbit insertion, station-

keeping and includes Reaction Control System (RCS) thrusters as external Attitude

Control System (ACS) actuators. The top level requirements for the satellite wet mass

(≤ 300 kg) and the mission lifetime (≥ 6 months) drive the sizing of the propulsion

subsystem. It will be sized such that the wet mass attains 300 kg, utilising any

unclaimed wet mass for additional expellant to extend the mission lifetime. The

propulsion subsystem restricts the mission lifetime via the expellant lifetime, i.e. the

duration for which expellants for station-keeping and RCS functions are available.

The design objective is to maximise the expellant lifetime by configuring a propulsion

subsystem that stores the maximum possible expellant mass. The expellant lifetime

is used by the systems engineer to determine the mission lifetime.

Sizing the propulsion subsystem is a recursive root-finding process, with the satellite

wet mass as the objective function. The independent variable is the stored expellant

mass at arrival in the mission orbit. With these two values, and a description of the

propulsion subsystem configuration, the propulsion subsystem is sized and its mass is

determined. Consequently the satellite wet mass is computed. This process is repeated

until the wet mass converges to the band [300−∆m, 300] kg, where ∆m = 1/10kg is

a mass increment specifying the accuracy of the calculation and also acts as the step

size during iterations.

4.2.1 Propulsion Subsystem Sizing

The required thrust levels and total impulse of the Main Engine (ME) and RCS

thrusters are small, thus pressure-feeding is acceptable to provide the necessary inlet

pressures. A trade-off study contrasting pump-fed and pressure-fed systems has not

been conducted. The mass of the propellant and oxidiser tanks are computed via

fitting a first degree polynomial to points for the tank mass and volume obtained

from readily available surface tension tanks manufactured by Astrium [14], because

sizing the fluid capture mechanisms is outside the scope of this preliminary study.

The functions obtained for Titanium MMH and MON tanks are mt = 34.35Vt +

8.387 (R2 = 0.898) and mt = 34.70Vt + 10.51 (R2 = 0.995) for spherical and

cylindrical (with domes) surface tension tanks respectively. Similarly for Titanium

surface tension Hydrazine tanks the mass is found to be mt = 79.25Vt + 1.52 (R2 =

0.780). Residual oxidisers and propellants are estimated to be five percent of the

stored mass.

For the regulated pressure-fed systems the required mass of pressurant is computed


according to [17] by:

m0 =plVl

RT0

(k

1− pp/p0

). (4.2.1)

Where m0 is the initial pressurant mass, pl and Vl are the gas pressure and volume

in the liquid tanks at their depletion respectively, R and k are the gas constant and

specific heat ratio of the pressurant, pp and p0 are the pressures in the pressurant tank

at depletion of the liquid and initially respectively and T0 is the initial temperature

in the tank.

This amount of pressurant provides the required inlet pressures for thrusters and

the ME at the end of life of the propulsion subsystem (i.e. at the depletion of the

tanks). A five percent margin is included in the pressurant mass for ullage and

residuals in the lines. The tanks for the gaseous pressurants (Helium or Nitrogen) are

sized for a given initial pressure as spherical shells with a constant thickness that is

determined via the ultimate tensile strength of Titanium, including a safety factor of

1.25. The initial pressure is 200 bar [15].

The propulsion subsystem plumbing (i.e. feed lines, valves, pressure regulators)

is accounted for by mplumbing = c mPS,dry with c ∈ [7, 17.5]%. The constant c is

modified to reflect the complexity of the propulsion subsystem, i.e. higher values

for bi-propellant systems than for dual-mode systems [18]. Propulsion subsystem

instrumentation and electronics is not accounted for explicitly, it is included in the

20% subsystem margin.

4.2.2 Propulsion Subsystem Configuration

The five different propulsion subsystem configurations under consideration are a mono-

propellant system (four MEs: Table 4.2.2 (f), RCS: Table 4.2.2 (a)), four mono-

propellant MEs (Table 4.2.2 (f)) and CG RCS (Table 4.2.2 (b)), a bi-propellant

system (ME: Table 4.2.2 (i), RCS: Table 4.2.2 (e)), a bi-propellant ME (Table 4.2.2

(i)) with CG RCS (Table 4.2.2 (b)) and a dual-mode system (bi-propellant ME:

Table 4.2.2 (l), mono-propellant RCS: Table 4.2.2 (a)). These configurations are

the conclusions of numerous trials on the basis of components the majority of which

is listed in Table 4.2.2. During development, procedures size each configuration si-

multaneously (using the same propulsion subsystem requirements as inputs) and the

resulting expellant lifetime is computed.

For the sizing of the propulsion subsystem, the most important system charac-

teristic is the dry mass of the satellite’s subsystems, as it determines (in combination

with the propulsion subsystem and structure mass) the mass allocation for expellants.


40 45 50 55 60 65 70 75 80 85 90−200

0

200

400

600

800

1000

1200

1400

1600

1800

Satellite dry mass excl. PS and structure [kg]

Expella

nt lif

etim

e [days]

Bi−propellant PS

Dual−mode PS

Mono−propellant PS

Bi−propellant ME, CG RCS thrusters

72.86 kg

272 days

Figure 4.2.1: Effect of the choice of propulsion system configuration on the missionlifetime via the subsystem dry mass (excluding the propulsion subsystem andstructure). The selected dual-mode configuration is annotated.

Figure 4.2.1 illustrates this behaviour for four different propulsion subsystem con-

figurations (the monopropellant system with cold gas RCS is omitted).It is found that

for the subsystem mass of 72.86 kg a dual-mode propulsion subsystem delivers the

highest expellant lifetime (272 days). Additionally, the use of mono-propellant RCS

thrusters in the dual-mode configuration reduces the subsystem complexity compared

to a bi-propellant system. The configuration comprising cold gas RCS thrusters,

although delivering a similar expellant lifetime, is unattractive as it does away with

having a shared propellant tank for the thrusters and the ME. A shared tank allows

the thrusters to use a propellant amount different from the design value, e.g. in the

case of having to re-purpose them for station-keeping in the event of an ME failure or

if the RCS propellant consumption is higher than expected. Therefore the dual-mode

propulsion subsystem con figuration is selected for the direct chemical transfer option.

The dual-mode propulsion subsystem is found to exhibit the highest expellant lifetime

and is thus the selected propulsion subsystem configuration. It is comprised of the

Northrop Grumman bi-propellant engine Table 4.2.2 (l) running on Hydrazine and

Nitrogen Tetroxide and 12 EADS Astrium mono-propellant Hydrazine thrusters Table

4.2.2 (a). Both share a common propellant tank. The propellant and oxidiser are


pressurised by a common Helium tank via separate lines and pressure regulators.

The impulse delivered by the ME during transfer follows from the expellant mass,

engine thrust and specific impulse as 3.24 × 105 Ns, i.e. only 19 % of the MEs total

impulse capability of 1.67×106 Ns. The remainder is available for station-keeping and

RCS functions. The RCS impulse delivered during transfer is negligible in contrast to

the 1.80× 105 Ns capability of a single RCS thruster. With the annual RCS impulse

requirement of 920 Ns it is clear that neither the lifetime of the ME nor the RCS

thrusters are constraining the mission lifetime.

A breakdown of the dry propulsion subsystem in terms of mass is given in Table

4.2.1.

Component Mass Allocation[kg] (%)

Oxidiser tank 9.74 27.6Propellant tank 10.05 28.5Pressurant tank 2.38 6.7RCS thrusters (12) 3.48 9.9ME 6.03 17.1Plumbing 3.17 9.0Pressurant 0.43 1.2Propulsion system dry 35.29 100.0(incl. pressurant)

Table 4.2.1: Mass breakdown of the dry dual-mode propulsion subsystem for thedirect chemical transfer option.


Manufacturer

Ref.

Designation

Thru

stI s

pPro

p.

Ox.

Mix.

Inlet

Mass

Acc

um.

ratio

pres.

burn

life

[N]

[s]

[bar]

[kg]

[hrs]

(a)

EADSAstrium

[13]

Mon

o-PropellantThruster

1220

N2H

4n/a

n/a

0.29

50(D

Cdual

seat

dual

solenoidvalve)

(b)

Moog

[16]

SolenoidActuated

Thruster

3.5

71.5

GN2

n/a

n/a

14.8

0.022

16666

58-118

(c)

EADSAstrium

[12]

Bi-PropellantThruster

10291

MMH

N2O

4,0.35

70(single-seat

valve)

MON

(d)

EADSAstrium

[13]

HydrazineThruster

20224

H2H

4n/a

n/a

0.395

10.5

(e)

EADSAstrium

[12]

Bi-PropellantThruster

22290

MMH

MON

0.65

(f)

NorthropGrumman

[11]

Mon

opropellantThruster

MRE-15

66228

N2H

4n/a

n/a

18.96

1.10

9.12

(g)

Ampac

ISP

[16]

MONARC-90

90235

N2H

4n/a

n/a

1.00

(h)

EADSAstrium

[10]

Bi-PropellantEngineS400-12

420

318

MMH

N2O

4,1.65

103.60

8.3

MON

(i)

EADSAstrium

[10]

Bi-PropellantEngineS400-15

425

321

MMH

N2O

4,1.65

104.30

12.8

MON,

(j)

Ampac

ISP

[16]

MONARC-445

445

235

N2H

4n/a

n/a

1.60

(k)

Northrop

[11]

Dual

ModeLiquid

Apogee

Engine

471.5

322

N2H

4N2O

41

14.13

4.763

6.71

Grumman

TR-308

(l)

Northrop

[11]

HighPerform

ance

Dual

ModeLiquid

556.0

330

N2H

4N2O

41.06

15.85

6.033

0.83

Grumman

Apogee

EngineTR-312-100YN

Tab

le4.2.2:

Selection

ofengines

andRCSthrustersusedfortheconfigu

ration

ofpropulsionsubsystem

s.


4.3 Power Subsystem

4.3.1 Requirements

To provide continuous power to the subsystems and payloads, their demands must

be accurately estimated. Since not all instruments are constantly operating at their

peak conditions, a power plan of the spacecraft subsystems must be compiled over one

average orbit. To accomplish this, a table is constructed containing all instruments

on-board in a row, and a time axis in a column in 5-minute interval over the course

of one orbit. Each cell in this table contains the power level from 0% to 100% of

the peak power, indicating the operational state of each instrument at a particular

time. This allows the calculation of the total instantaneous power requirement of the

spacecraft over the course of one orbit, plotted in Figure 4.3.1.

The average figure of the instantaneous power requirement was then found to be

115.2 W. To account for information lost by using discreet time, a contingency of 10%

is applied to this value. This is also to slightly oversize the power subsystem to ensure

the battery being fully charged after each sunlit phase. This, Therefore, gives the

subsystem power requirement (PSR) of 126.7 W. To carry on with sizing the batteries

and solar arrays, it is assumed that the spacecraft steadily consumes this amount of

power over the course of one orbit.

Figure 4.3.1: Power profile outlining the power demand of an average orbit and thewosrt-case sunlight condition.

4.3.2 PCDU

To provide the payload with stable electric power, the voltage and current fed into the

payload and subsystem must be regulated. To do this job, a Power Conditioning and

Distributing Unit (PCDU) is used. The schematics of power subsystem is described

in Figure 4.3.2.

4.3 Power Subsystem 23

Power produced by the array is regulated by a PCDU. Since the subsystem power

requirement is 126.7 W, and the solar array, as described below, only produces 180

W, Small Satellite Power System produced by SSTL can be used, which is scalable

upto 1.6 kW. [19]

This process occurs with some finite DC-DC efficiency, (µPCDU). Since the product

description for Small Satellite Power System did not include this efficiency, a value is

taken from a similar PCDU by Thales Alenia, which is 94%. [21]

25% of subsystem mass has been allocated for cables.

Figure 4.3.2: A schematic on the power subsystem. During daylight the solar arraypowers the payload and subsystems and charge the batteries, via a PCDU. Duringnight time, the batteries power the load via the PCDU.

4.3.3 Battery Sizing

It has been found that the power subsystems require (PSR) is 126.7 W. The battery

must be sized to enable satellite operation during all encountered eclipses. To do this,

the longest possible eclipse time is used. Given AoP of 320.6◦, the maximum eclipse

time (Teclipse) was found to be 3023 seconds. The minimum energy capacity of the

on-board battery Emin is therefore given by:

Emin =PSR × Teclipse

DoD × µPCDU

(4.3.1)

It was decided that VES 180 Li-ion batteries produced by SaftBatteries would be

used for their high specific energy of 175 Wh/kg. This battery is capable of achieving

60,000 cycles at 20% DoD. By linearly scaling this figure, it is found that even if the

battery were to be 100% discharged, it would be capable of achieving 12,000 cycles.

Since in the 10-month life the battery is going to discharge about 1350 times, the

battery should not wear by the cycles at 100% DoD. To ensure that the battery does

not completely discharge in an eclipse, however, DoD of 80% has been chosen. [22]


Manufacturer Saft Batteries

Name VES 180

Dimensions

[mm]

53 ϕ x 250

Discharge

Voltage

3.6 V

Capacity 50 Ah

Mass 1.11 kg

Max Discharge

Current

100 A

Max cycles at

20% DoD

60,000

Figure 4.3.3: Specifications of VES180 Li-ion battery produced by SaftBatteries

Using these figures in Equation (4.3.1),

it was deduced that the battery must

at least contain 154 Wh. Since one

VES 180 battery features 180 Wh, one

battery is enough to power the mission.

The spacecraft peak power requirement

is also within the maximum discharge

power of 360 W, and since charging

time is longer than discharge time, the

battery charging current will not exceed

the discharge current.

Hence, one of this battery is chosen for

the spacecraft.

4.3.4 Solar Array Sizing

During daylight, the solar array must

provide for both PSR and charging the

battery. This is the minimum power that would allow the spacecraft to nominally

operate, and hence this power is used to define the End-of-Life power (PEOL). PEOL

is then given by:

PEOL =PSR

µPCDU

+DoD × Emin

µBattery × Tdaylight

(4.3.2)

where µBattery is an efficiency figure for charging and discharging the battery. For this

study, this efficiency is set to 1.

It is shown that minimum power required by the solar arrays is 162.2 W.

Over the course of the mission, the solar arrays will be subject to damaging charged

particle fluences. Data has been collected on such fluences in the trans-lunar space,

and a study has shown that most of such particles are trapped in the Van Allen

radiation belts. [25]

Using particle fluences measured in the radiation belts, and the product speci-

fication by emcore, damage done during the 21-day commissioning sequence in GTO

has been found. Using coverglass of 152 µm thickness, it was shown that during com-

missioning, SA degradation P/PBOL is 90%. [24, 20] Degradation at lunar radiation

environment was estimated to be 0.1% per year.

Using these figures, Beginning-of-Life power requirement is calculated:

PBOL =PEOL

Dtransfer × (1−Dlunar)Tmission

1year

(4.3.3)


where Dtransfer is a one-off degradation during transfer, and Dlunar is the degradation

per unit-time spent in lunar orbit.

Using Tmission = 10 months, this has produced BOL power requirement of 181.8 W.

As shown below, the solar arrays will be body-mounted on five faces. This means

that at any given moment, more than one face will be illuminated. To ensure that

PBOL is at all times produced when the satellite is illuminated, it has been decided

each face must be capable of producing PBOL when illuminated perpendicular to the

sunlight, whih is when the total projection-area of all solar arrays combined would be

at its minimum.

ZTJ PV Cell produced by emcore has an efficiency of 29.5%. [20] Using these cells,

the Sun-projection area required to produce PBOL is:

ASA =PBOL

µcell × Φsun

(4.3.4)

Here, packing efficiency is not considered, as the number of cells required are to be

found. Using Φsun = 1,350 W/m2, this has produced sun-projection area of 0.46 m2.

4.3.5 Solar Array Configuration

It had been recognized that the deployed arrays would introduce a single point of

failure, and would strain the AOCS, as it would increase the moment of inertia of

the spacecraft by a factor of as high as 2. As it had been found when sizing the

power subsystem for the low thrust option, the increased moment of inertia meant

that a heavier set of reaction wheels (by a factor of 3) had to be used, as explained

in Section 6.6.3. The mass-increase in choosing the heavier wheels would not allow

the spacecraft to carry enough fuel, consequently it would force the elimination of

secondary payloads. The low thrust option was able to accommodate these wheels

due to the smaller fuel mass required, however, this would not be the case for chemical

options. It has therefore been decided that this configuration will only be employed

when the sun-projection area of the array exceeds the area of one side of the spacecraft.

Since the satellite is cubic with 1 m side-length, as shown in Figures 4.8.1 and 4.8.2,

this limit would be 1 m2. This situation has been avoided, hence body-mounted

solution is employed.

The solar arrays must be placed so that, when in daylight, are producing enough

power to power the subsystem and charge the batteries. Placing the arrays on all six

sides of the spacecraft would produce a fully power-safe spacecraft, however, it would

strain the thermal control subsystem, as GaAs arrays have high emissivity, rendering

the spacecraft very cold during eclipse, as shown in Section 4.7. It was found that

placing five arrays would also provide power-safety, as well as reduce the strain on the

thermal control subsystem. Hence solar arrays are placed on five faces of the satellite.


4.3.6 Power Profile

Having selected the subsystem components, the power profile over one orbit is simulated.

Power produced by the arrays and power demanded is compared, such that the battery

is charged when the power is in surplus, and discharged when the power is in deficit.

The result is plotted in Figure 4.3.1 The dip in the SA power production is due to

an eclipse; the longest eclipse at aposelene is used to produce this graph. After half

a sideral lunar rotational period, the shortest eclipse is observed at the opposite side

of the moon.

4.4 On-Board Data Handling Subsystem (OBDH)

The On Board Data Handling (OBDH) subsystem takes care of spacecraft command

and control by means of a micro-controller. The actions to be performed are defined

both by a software stored in the system and by command up-linked from ground

stations.

The spacecraft has to work in three different phases: Geostationary Transfer Orbit

(GTO) parking, transfer, and lunar orbit. Among these, the third is the more com-

putational demanding since the science data processing will be added to all the usual

tasks (e.g., data housekeeping, solar arrays and antennas pointing). Therefore, the

On-Board Computer (OBC) is sized to meet lunar orbit computation demands.

4.4.1 Overview

The most important system duties are:

• Science data collection and processing.

• Monitoring of the spacecraft health via sensors data.

• Act as communication hub between subsystems, and between the ground station

and the spacecraft.

A system capable of handling these data can be designed as depicted in Figure 4.4.1.

The OBC receives and decodes commands which are then directed toward the payloads

and other subsystems by operating a digital to analog (D/A) conversion where needed.

The subsystems operate the required operations and return science and telemetry

data. These are marked with a unique time-stamp using the internal clock and then

processed by the on board software. Accordingly to ground control directives, storage

and down-link capabilities, the collected data are selected, usually giving priority to

science ones. In order to optimise the down-link process, different types of data are

merged and possibly compressed through the multiplexing step. Eventually all the

compressed packages are stored until the next available communication window when

the information are encoded and sent to the ground station.

4.4 On-Board Data Handling Subsystem (OBDH) 27

..

...Earth’sGroundStations

. . . ...Moon

.

. ...Receiver ...Decode

. . . . ...Payload

. . . ...Analog/DigitalConverter

. . . . ...Subsystems

. . . ...Clock

. ...DownlinkCapabilities

...DataSelection

. . ...Multiplex ...Storage

. ...Transmitter ...Encode

.

commands

.science

.

telemetry

.

Communications

.

On-Board Data Handling

.

Lunar Orbiter

Figure 4.4.1: OBDH system scheme showing data flows between subsystems as wellas between the spacecraft and both ground stations and the Moon’s environment.

4.4.2 Requirements

The OBC should be able to handle and store all the information gathered in between

at least two down-link sessions i.e., one lunar orbit. In Table 4.4.1 the types of data

and their size are reported.

For each orbit at least 10 pictures per camera are taken.

4.4.3 Components

This section focuses on OBDH components and their selection. In Figure 4.4.2 the

ensemble of system parts is organised to show their relationships.

The OBC has been assembled using only off-the-shelf components. Therefore,

each part described in the following sections comes from the catalogues of actual

manufacturers. Although several catalogues have been searched for each component,

only the chosen component is reported, along with a brief description.


Source SizeHigh Res. Camera 100MB/pictureLow Res. Camera 44.352MB/pictureDust Detector 150KB/dayRadiation Detector 400KB/daySpectrometer 36Mb/orbitTelemetry 300 b/s

Table 4.4.1: Data sources and sizes. The raw file size is taken into consideration.

..

. ..COP . . . . . . ..DigitalI/O 1

..Real-TimeClock

. . . . . ..Driver

..PROM . . . ..CPU

..SRAM . .. . . . ..Driver

. ..EDAC . . . . . . ..DigitalI/O n

..FLASH . ..

..AnalogInput 1

.. . . . . .. .. ..AnalogOutput 1

. . ..Muxer ..A/D . ..D/A ..Demuxer

..AnalogInput n

.. . . . . . .. ..AnalogOutput n

Figure 4.4.2: OBDH scheme in details. The most important links between componentsare shown.

The processor has been selected paying attention to its performance, available

cache, power consumption, radiation resistance, and heritage. In addition, two fun-

damental requirements were the space operation qualification and a 32-bit architecture.

The picked micro-processor is the RAD750 produced by BAE Systems [35, 37].

The decision has been driven primarily by its computational power (400MIPS) and

in second instance by its heritage. Indeed, it has been employed in missions like the

Mars Science Laboratory (MSL) one, where it proved high reliability and the capacity

of meeting high computational demands thanks to its high clock frequency (200MHz),

which is supported by 1MB cache.

It is not unusual that for a specific micro-processor architecture a particular

operating system is written. Indeed, the RAD750 is often used in combination with


the Wind River VxWorks Real Time Operating System (RTOS) [34]. This Operative

System (OS) has been used to coordinate the landing operations for the MSL, an

operation which required high accuracy and reliability. VxWorks system requirements

are reported in Table 4.4.2.

Minimum [KB] Recommended [KB]RAM 1000 4000ROM 128 2000NVRAM 512 512

Table 4.4.2: Wind River VxWorks RTOS system requirements.

The Read Only Memory (ROM) will be used to store the operating system bootable

image, while the Non-Volatile RAM (NVRAM) will contain boot line information

needed in case of malfunctioning. The Random Access Memory (RAM) requirement

is relative to the only OS, hence the subsystem will be likely to require a bigger

memory to run additional software.

In the view of designing the system on qualified circuitry, it is common to use

Single-Board Computers (SBCs) i.e., small size (generally 320mm ×170mm ×55mm)

boards containing all the needed electronics. Generally, these computers are equipped

with a particular micro-processor and a set of memories. Moreover, useful devices

such as a Clock and a Computer Operating Properly (COP) timer are embedded on

the board. The former is responsible for the synchronisation of time between all the

subsystems; it is needed in order to schedule all the tasks and to provide a unique time-

stamp to be attached to the down-linked data. The latter consists basically in a timer

which is automatically reset by the on-board software with regular basis. Whenever

the software fails i.e., the system stops working, the timer reaches a threshold which

has the effect of restarting the entire system. Eventually, bit errors can be prevented

with encoding/decoding processes as well as using Error Detection And Correction

(EDAC) units in between memories and processor.

The single-board option does not prevent the designer to expand the computer

capabilities, indeed the manufacturers themselves usually provide the possibility to

customise their devices.

The chosen SBC is the BAE System 3U cPCI SBC [36]. Its strengths reside in

the small dimensions (320mm ×170mm ×55mm) and light weight (549 g), as well

as the massive amount of RAM available (128MB). Indeed, in the view of subsystem

redundancy, there will be two identical OBC in the spacecraft, hence a small and

light circuitry is certainly preferable. It also includes both the timer and the COP.


The only drawback is the small ROM size (256KB Start Up ROM (SUROM)), which

is not big enough to contain the OS bootable image. Hence, a dedicated external

memory module has to be added.

Accordingly with the VxWorks’s recommended system requirements (Table 4.4.2),

the total non-volatile storage size needed is 2.512MB. Given the SUROM of 256KB

actually present in the SBC, the additional memory unit has to be at least 2.256MB

in size.

Non-Volatile RAMs (NVRAMs) have the advantages of both RAM and ROM i.e.,

they are rewritable and can hold data for theoretically infinite time. Hence, to increase

the ROM size, the BAE System C-RAM 4M RAD NVRAM [38] has been added to

the SBC. It provides 4MB of space to be used for the OS needs. Furthermore, by

choosing a component made by the same manufacturer of the SBC, the compatibility

between parts is ensured. Indeed, BAE Systems allows to customise its radiation

hardened single board computers in order to meet the the subsystem engineer needs.

Storage size requirements can be computed taking into account the values in Table

4.4.1 and the orbital period of 19289.33 s. The primary science data size is 1440MB.

This value is further increased of the 2% in order to consider the overhead information,

this bring the total pictures size to 1469.38MB. The secondary payloads contribute

with 40.72MB per orbit, where the 10% overhead has been considered. The telemetry

data produced during one orbit are 6.43MB, where the 10% has been considered.

Eventually the total amount of data to be handled is 1516.53MB.

The Space Micro RAD NAND Flash Module can be mounted directly on the SBC

[32], and allows to store up to 8GB of data. The Flash unit completes the OBC

memory set.

The internal physical network joining the OBC to the subsystems and payload is

developed by using the SpaceWire standard [30]. It allows to share data between

two devices at up to 200Mbit/s in both ways. This value ensures that there will not

be delays in the data transmissions, leaving process time issues to the OBC compu-

tational capacities.

In order to use SpaceWire standard, the OBC needs a particular interface card, a

router with as many ports as the number of subsystems, and cabling. BAE System

provides, along with the chosen 3U cPCI SBC, the 3U cPCI SpaceWire interface card

which is equipped with four on-board ports. It weights 0.5 kg, resists up to 100 krad,

and its dimensions are 100mm×160mm×30mm.


..

..Communications .. ..SBC .. ..Camera 1

..AOCS .. ..SW cPCI . ..Camera 2

..Power .. . .. ..Dust Detector

..Propulsion . ..SW Router . ..RadiationDetector

..Thermal . . . ..Spectrometer

.

1

.

2

.

1

.

3

.

4

.

2

.

3

.

4

.

5

.

6

.

7

.

8

.

OBC module

Figure 4.4.3: Spacecraft SpaceWire internal network.

By considering the total number of subsystems and science instruments, it is clear

that there is the need for more than four ports. This is why a SpaceWire router has

been looked for. The router itself consists in only a small chip, indeed the assembly of

the network facility is left to the spacecraft engineers. Being this not a feasible option

an off-the-shelf router has been found. The SPA SpaceWire Network Router [33],

manufactured by Design Net Engineering LLC, is configurable with up to 16 ports,

despite no more than 8 ports would be needed (Figure 4.4.3). The router size is given

as 127mm×178mm×50mm [33]. By using this value a rough mass approximation can

be done. Since it takes up space of about 1.5 OBCs, then its weight is approximated

to 1.5 kg in order to take into account its shielding.

Cabling mass has been considered as the 20% of the total OBDH mass.

To complete the connection between the different devices, only an analog interface

is needed. In particular, a signal converter analog/digital/analog is manufactured by

Honeywell in form of on-board chip [31], and can be added to the SBC in the same

way the FLASH unit has been integrated.

The hardware has now been selected completely, the final configuration and re-

dundancy computations are reported in Table 4.4.3. In particular, to take into account

the redundancy, the entire system has been duplicated.

4.4.4 Software

The development of the spacecraft software has not been considered. Indeed this

process may take several years of code optimisation. Also, one of the main OBC

task, the attitude regulation, is performed by a separate computational unit which

is integrated into the Attitude and Orbital Control System (AOCS). Hence, the only

assumptions done are about the compression of science and telemetry data for com-

munication purposes.


Images are compressed initially by using the JPEG format. Besides the JPEG

compression is not loss-less, a small compression ratio, say 1 : 10, introduces a small

amount of noise in the picture, nearly invisible to the human eye. Therefore, all the

images taken are reduce to one tenth of the raw size. Subsequently, a loss-less com-

pression algorithm is used to apply a further compression ratio of 1 : 1.76 to all the

collected data [29]. Eventually the total data size is reduced to 110.27MB/orbit. This

value represents the output given to the Communications subsystem to estimate the

down-link requirements.

4.4.5 Considerations for Direct Chemical

The direct chemical transfer option does not require any additional component aside

from those reported in the general OBDH configuration.

In order to assure a proper radiation shielding, the results obtained by the spacecraft

Chandrayaan-1 during its journey to the Moon have been used [27]. Initially a time

period of two weeks in GTO has been assumed. The continuous passage through

the Radiation belts will provide a total amount of 1.0848 krad. The short five days

transfer will add 0.144 rad. Eventually, during the six month of orbital mission the

amount of radiation is estimated to be of 669.071 rad. The total amount of radiation

for the entire lifetime is 1.754 krad.

Single Unit RedundancyPerformance 400MIPS @ 200MHzMemory 1MB cache

128MB SDRAM256KB SUROM4MB NVRAM8GB FLASH

Mass 2.6 kg 5.2 kgDimensions 127×178×110mm 127×178×220mmPeak Power 22.8WRadiation 100 kradOperating Tem-perature Range

−55◦C +70◦C

Cabling 20% subsystem massTotal mass 2.76 kg 5.52 kg

Table 4.4.3: OBDH configuration. Redundancy is taken into account by duplicatingthe whole subsystem.

4.5 Communication Subsystem 33

4.5 Communication Subsystem

The communications subsystem will carry out a number of key functions.

• Transmit housekeeping data to flight controllers.

• Receive instructions from flight controllers.

• Transmit photographs.

• Transmit scientific data.

4.5.1 Link Budget

Evaluation of the link budget is performed by selecting components and planningoperations such that required data rates can be met while minimising mass, andpower demands to an acceptable level.The link budget is calculated using Equation 4.5.1.

10 log10

(C

N0

)= 10 log10(PTGT ) + 10 log10

(GR

TR

)− 20 log10

(4πρ

λ

)− 10 log10 (LA) − 10 log10 (k) (4.5.1)

Where

CN0

carrier power to noise densityratio

PT transmitter RF power

GT transmitter gain TR receiver equivalent noise tem-perature

ρ slant range from transmitter toreceiver

λ wavelength

LA atmospheric loss k Boltzmann’s constant

The following passages will progressively define each of the terms in this equation.

The CDE spreadsheets are used to calculate a solution for each mission phase, where

the final output in each case was the RF power required for the transmitted signal.

By considering communication windows imposed by the mission orbit and attitude

selection, the quantity of scientific data gathered was optimised to satisfy mission

requirements while keeping this power within the capabilities of selected components.

4.5.2 Link Quality

To define the CN0

, a typical bit error rate of 10−5 was chosen. This gives Eb

N0of 10dB

[39]. Carrier power to noise density ratio is obtained by multiplying this density by

bit rate (Rb) according to Equation 4.5.2.

C

N0

=Eb

N0

Rb. (4.5.2)


Since the concurrent design approach meant that bit rate would vary during design,

this equation was used in the concurrent design environment spreadsheets to calculateCN0

, while bit error rate was assumed constant.

4.5.3 Ground Stations

ESA Tracking Stations (ESTRACK) will provide sufficient coverage and data rates

at typical frequency bands and bandwidths. Using STK, ESTRACK coverage was

investigated and it was found that by using tracking stations in Kiruna, Kourou, and

Perth, the spacecraft would be in almost continuous visibility (greater than 99%)

provided the Moon is not blocking the line of sight.

All subsequent investigation into the communications coverage assumes these tracking

stations are to be used, their basic information is tabulated in Table 4.5.1

Kiruna-1 Kourou-1 Perth-1

Dish diameter 15 m 15 m 15 mS-band RX band 2200 - 2300 MHz 2200 - 2300 MHz 2200 - 2300 MHz

S-band G/T 27.7 dB/K 29.1 dB/K 27.5 dB/KX-band RX band 8025 - 8500 MHz 8025 - 8500 MHz 8025 - 8500 MHz

X-band G/T 36.9 dB/K 41 dB/K 37.5 dB/KData rate < 100 Mbps 2 Mbps 2 Mbps

Table 4.5.1: Performance characteristics for the Kiruna-1, Kourou-1 and Perth-1terminals [40]

4.5.4 Carrier Frequencies

High gain transmissions will use X Band. This band was selected for high gain trans-

missions because it has heritage with high data rate transmission, and is compatible

with the ESTRACK network. S-Band was chosen for low gain communications for

the same reasons.

It is clear from the data on the selected tracking stations and the data rates calculated

for the OBDH subsystem that the bandwidth required by the spacecraft for high and

low gain transmissions can be accommodated by the selected ground stations at X

and S bands.

To carry out link budget calculations in the CDE, arbitrary carrier frequencies within

the relevant bands given above have been assumed. These are 8450MHz for high gain

communications and 2250MHz for low gain communications.


4.5.5 Losses

Both free space (LFS) and atmospheric losses (LA) in the beam must be considered.

Free space loss is given by Equation 4.5.3.

LFS =

(4πρ

λ

)2

, (4.5.3)

Assuming a maximum distance to be the apogee radius of the Moonfs orbit, 405400km,

we find that (LFS) is 1.43×10−11 for high gain transmissions and 3.82×10−10 for low

gain transmissions.

Using frequency bands below 10GHz, spacecraft communications will not suffer sig-

nificant clear air attenuation [41]. Assuming an atmosphere of dry air and water

vapour, zenith attenuation is approximately -0.04dB [42] for spacecraft transmission.

As well as attenuation due to dry air, some loss would be caused by precipitation.

However in this case it is not a significant factor. If it were to be considered, the

level of attenuation due to precipitation may be estimated using rainfall statistics and

the method described in various documents by the International Telecommunications

Union [39]. This method consists of sizing the communications system to handle at-

tenuation levels that will be not be exceed for some acceptable percentage of the time,

eg. 99.9%.

Over the distances concerned, free space losses occur at a much greater order of

magnitude than atmospheric losses. Atmospheric losses are around 2dB at 11GHz for

a link reliability of 99.9% [39], while free space losses are 223dB. For this reason it is

deemed an acceptable simplification not to calculate a value for attenuation due to

precipitation, and instead use only dry air attenuation in calculating the link budget.

4.5.6 Component Selection

Given the high free space loss compared with missions to LEO, the transmitter will

need to have sufficiently high power and gain in order to reach the required data rate.

After reviewing available hardware options, the following communications architecture

was selected.

• 1x X-band steerable high gain antenna.

• 1x X-band transmitter.

• 6x S-band low gain antennas.

• 2x S-band transmitters.

• 2x S-band receivers.


Figure 4.5.1: Zenith attenuation of signal vs frequency [42]


All of these components are available from SSTL, where both the X-band components

and S-band components have flight heritage together. The specific hardware and their

key data are given in 4.5.2, 4.5.3, 4.5.4, 4.5.5, 4.5.6.

Mass 3 kg for 15dBiC, 3.3 kg for 18dBiC

Power1.3 W static

3.9 W dynamicFrequency range 8000 - 8500 MHz

Gain 15 dB or 18 dBSlew rate < 20◦/s

Azimuth Range +/- 270◦

Elevation range +/- 110◦for 15dBiC, +/-80◦for 18dBiC

Table 4.5.2: Key data for the SSTL X-band high gain antenna pointing mechanism[43].

Mass 4 kg

Power demand65 W at 5 W RF

120 W at 12 W RFFrequency range 8025-8400 MHz

Data rate 10.0 - 500.0 Mbps

Table 4.5.3: Key data for the XTx400 X-band transmitter [44].

Mass 3 kg for 15dBiC, 3.3 kg for 18dBiCFrequency range 2000 - 2500 MHz

Gain -5 dB at 90◦off boresight

Table 4.5.4: Key data for the SSTL S-band patch antenna [45].

Mass 1.8 kgPower demand at 4 W RF < 38 W

Frequency range 2200 - 2250 MHz

Table 4.5.5: Key data for the S-band downlink transmitter [46].

Mass 1.3 kgPower demand 1.5 W

Frequency range 2025 - 2100 MHzData rate 9.6 kbps or 19.2 kbps

Table 4.5.6: Key data for the S-band uplink receiver [46].

These pieces of hardware were selected as they would roughly satisfy RF powers

calculated using the link budget spreadsheets in the CDE derived from data rates


demanded by the OBDH subsystem. Their data was input and OBDH figures were

adjusted to keep the required RF power within the limits given in these tables.

The system will employ redundancy in S-band communications, but not in X-

band. This a mass saving measure; the need for redundancy was traded off against

mission life since some of the mass saved can be used for station keeping fuel, allowing

longer orbit maintenance and therefore prolonging the mission. It is possible that the

lack of redundancy for high gain communications could result in a severely reduced

transmission capability in the event of failure of the X-band antenna or transmitter.

In the case of the Galileo probe, when the high gain antenna failed to deploy the

mission controllers were forced to use low gain antennas to receive all scientific data.

A similar contingency is examined later in this chapter.

4.5.7 Investigating Communications Windows

Using Kiruna, Kourou and Perth as ground stations, STK provides information on

typical transmission windows while in the final mission orbit around the Moon.

Most access windows last more than 5000 seconds. However there were also many

windows closer to 3000 seconds. Transmission will thus occur once per orbit, and a

typical worst case transmission window is assumed to be 2700 seconds (45 minutes).

Due to the elevation constraint for the X-band antenna at maximum gain, see Table

4, and the pointing requirements of the instruments, the communications subsystem

places a demand on spacecraft attitude control. Twice each month, when the plane

of the mission orbit is almost normal to the vector to the Earth, the limit in antenna

elevation requires the spacecraft to slew for each transmission. This repointing is

necessary as long as the orientation of Earth relative to the orbit places the ground

stations outside the field of view of the antenna, using STK it was found this slew

is required for 8 orbits during these periods. Communications slewing will therefore

occur for a total of 16 orbits per lunar month, once during each of these orbits. The

degree of slewing will be 15◦to ensure that the antenna can point directly at the

ground station being used.

Due to the almost constant ground station visibility during transfer, transmission

windows were not investigated for this phase of the mission. It was also assumed that

since no scientific data is gathered during transfer, only low gain communication would

be used. By placing an S-band patch antenna on each spacecraft face the need for

communications slewing is eliminated during transfer. The small mass contribution

of multiple S-band patch antennas (80g each) was accepted to reduce the fuel mass

that would be required if the spacecraft were to be slewed during this mission phase.


Figure 4.5.2: STK model of satellite communications. Cone shows field of view

4.5.8 Contingencies

Since the high gain capability of the communications subsystem will have no re-

dundancy, its possible failure, and the feasibility of using the low gain antennas to

transmit scientific data to earth, has been examined.

An alternative link budget was produced. The budget used for normal operation

takes data rates as an input and given information on the transmitter, receiver, and

some losses it will calculate the required RF power. The contingency link budget

reverses this process. It assumes that the low gain antenna will be running at full RF

power (4W) and calculates a maximum bit rate. This was found to be 36.45 kbps. This

rate was multiplied by the transmission window length to give a maximum quantity

of down linked data. How this quantity is divided between housekeeping and the

different types of scientific data is discussed further in section 4.4.


4.5.9 Final Results

The following table shows the final set of data rates to be used in the case of the direct

chemical transfer method. These values were calculated using the CDE spreadsheets

and are considered to be the final output for the communications segment.

Mission phase Data rate

Transfer orbit 300 bpsMission orbit (low gain) 1.37 kbpsMission orbit (high gain) 256.02 kbpsContingency data rate incase of high gain antennafailure

36.45 kbps

Table 4.5.7: Final key values for the direct chemical option.

4.6 Attitude Determination and Control Subsystem (ADCS)

The ADCS system will satisfy orbiter pointing requirements and steady the payload

so that science goals can be accomplished.

4.6.1 Control Modes

Firstly the control modes of the mission must be defined. Identifying these modes

will provide the requirements of the ADCS.

1. Initial Parking Orbit - The period between the orbiter being released from Ariane

5 into it’s initial Earth orbit and the beginning of the lunar transfer phase.

This will require ADCS control for testing equipment. These requirements are

negligible in comparison to the rest of the mission and so shall not be considered.

2. Transfer Slew - During this mode the Thruster will perform three burns; an

initial burn, a mid-way burn and a retro-burn. To ensure accuracy the orbiter

may need to be re-orientated.

3. Lunar Orbit Insertion and Acquisition - Initial determination of attitude and

stabilisation of vehicle upon arrival to the Moon. This mode may also be used

to recover from potential power upsets or emergencies. In order to account for

this, a safety factor will be included.

4. Nadir Data Collection - This mode occurs when the primary camera is nadir

pointing. During this mode the ADCS must provide a stable platform from

which pictures can be taken.

4.6 Attitude Determination and Control Subsystem (ADCS) 41

5. Off-Nadir Data Collection Slew - To achieve the orbiter’s science goals, it will

need to perform small slewing movements in order to image areas of interest

that are not covered by it’s ground track.

6. Communications Slew - The orbiter will have to slew back and forth from it’s

regular data collection mode in order to facilitate communications with the

Earth.

7. Eclipse Observation Slew - The orbiter will have to perform several large slews

to observe both the Moon and the Earth during the eclipse phase outlined in

the science goals.

8. Contingency Mode - A safety setting that can be implemented in an emergency.

Requirements for the four slewing modes are outlined in Table 4.6.1. For suitable

payload performance, the ADCS must fulfil the requirements given in Table 4.6.2 for

attitude determination and control whilst imaging. These requirements are applicable

to the data collection modes, and are driven by the primary payload.

Control Mode Slew Angle/◦ Slew Time/s Frequency Rate/◦s−1

Communication 15 60 0.54 per day 0.250Transfer 180 90 3 total 1.500Eclipse Observation 180 60 5 per year 2.000Off-Nadir Collection 5 60 10 per day 0.083

Table 4.6.1: Requirements for the four slewing modes - Mode 2, 5, 6 and 7.

Parameter RequirementAccuracy of Determination and Control 20 arcsecsRange Over Which Accuracy is Met ±180◦

Maximum Allowable Jitter 0.1◦/minDrift Allowance 1◦/hourSettling Time 1 min

Table 4.6.2: Camera Determination and Control Requirements

4.6.2 Selection of Attitude Control Method

In order to provide three-axis accuracy of determination and control to an accuracy of

20 arc seconds (0.00556◦), either a zero momentum Reaction Wheel Assembly (RWA)

consisting of three wheels or Control Moment Gyroscopes (CMGs) could be used.[48]

Other types of torquer simply wouldn’t be able to provide the same level of precision.

CMGs have historically been used for satellites greater than several tonnes such as the


ISS.[52] There has been a recent push to miniaturise CMG technology because they

offer a high mass and power-to-torque efficiency. However, currently the smallest,

reliable, off-the-shelf products such as Astrium’s CMG15-45S and Honeywell’s M50

CMG are designed to provide pointing requirements for satellites of roughly 1000 kg

and so would still be oversized for LEWIS.[49][50] Some CMG prototypes exist for

satellites nearer 300 kg, such as the University of Surrey’s SGCMG, but they are still

in development and so can’t be considered for this mission.[51] Therefore only a zero

momentum three wheel reaction wheel assembly can be considered for attitude control.

4.6.3 Quantifying the Disturbance Environment

This transfer method only requires us to quantify the disturbance environment around

the Moon. The greatest disturbances will be due to solar radiation pressure and

gravity gradient effects caused by the Moon. The orbiter will also be affected by the

gravity gradient effects caused by the Earth and the Earth’s magnetic field, however

those effects are negligible once in lunar orbit.

The worst case disturbance torques during lunar orbit orbit must be estimated. The

torque disturbance due to the gravity gradient effect on the orbiter, Tg is calculated

using,

Tg =3µ

2R3|K| sin(2θ) (4.6.1)

where R is the orbit radius (m), θ is the maximum deviation of the z-axis from the

local vertical, (to simulate the worst case scenario this will be fixed as 45◦) and K is

the greatest difference between the moments of inertia about z, y and x axes in kgm2

after transfer. The orbit radius will change over time, but the rest of the elements of

the equation are known, giving,

Tg =3× 4.90× 1012

2R3|3.31| sin(90◦). (4.6.2)

This equation can now be used to find the average disturbance torque due to the

gravity gradient effect of the moon yielding,

Tg = 8.40× 10−7Nm. (4.6.3)

The disturbance torque due to solar radiation pressure, Ts can be quantified by using,

Ts =Fs

cAs(1 + q) cos(I)(cps − cg) (4.6.4)

where Fs is the solar constant, 1367W/m2, c is the speed of light (3× 108m/s), As is

the largest surface area plane (0.5m2 for this transfer method), cps is the location of

solar pressure, cg is the centre of gravity, q is a reflectance factor (this is assumed to


be 0.5) and I is the angle of incidence of the sun (assumed to be 0◦ for the worst case

value). A value of 0.1 m is used for the difference between cg and cps for preliminary

design. This should account for changes in the geometry or for the selection of ad-

ditional payload. Therefore we get the result,

Ts =1367

3× 108× 0.5(1 + 0.5) cos(0◦)(0.1) = 6.83× 10−7Nm (4.6.5)

So the average total disturbance torque experience during lunar orbit, TL is,

TL = Ts + Tg = 1.52× 10−6Nm. (4.6.6)

4.6.4 Selection and Sizing of ADCS Hardware

To provide the camera requirement of an attitude determination accuracy of 20 arcsec

(0.00556 ◦) a Star Sensor must be used. The most suitable choice is the RIGEL-L star

tracker designed by SSTL. It exceeds requirements with a determination accuracy of 3

arcsec for less mass and power than comparable hardware. Using a 40◦ sun exclusion

baffle the star tracker will have a 40◦ exclusion angle to the sun and a 29◦ exclusion

to the Earth. In order to provide continuous attitude determination and to allow

for redundancy SSTL suggest the use of three CHUs (Camera Head Units) with one

DPU (Data Processing Unit) in the configuration shown in Figure 4.6.1. This can then

be mounted onto the face of the orbiter that points zenith during nadir data collection.

Figure 4.6.1: SSTL’s suggested configuration for RIGEL-L Start Tracker

The Star Tracker provides the required accuracy but LEWIS must be stabilised to

some extent before the tracker can begin to determine its attitude. Therefore there

is a need for a Sun sensor system which will be used during initial attitude deter-

mination, failure recovery and to aid attitude determination during slews. A suitable


option would be the space tested 2-Axis DMC Sun Sensor, also designed by SSTL.

With a field of view (FOV) of ±50◦, a sensor will be positioned on each face to provide

continuous attitude determination and allow for redundancy. These sensors have an

accuracy of 1.0◦, giving enough stability for the star tracker to take over. Additional

key specifications for both types of sensor are given in Table 4.6.3.

To provide 3-axis stabilisation control to a 20 arcsec degree of accuracy a reaction

wheel assembly must be used. Typically satellites of around 300 kg with body mounted

arrays use reaction wheels with an angular momentum capacity of 1 Nms so for initial

sizing, the chosen reaction wheel for this transfer method is the MW-1000 Reaction

Wheel developed by Space Quest Ltd, which satisfies this 1 Nms requirement. If

three-axis, zero bias, control with redundancy is required then there is a need for four

reaction wheels arranged as in Figure 4.6.2. The key specifications for the RWA are

given in Table 4.6.3.

Figure 4.6.2: Three-axis zero bias (redundant tetrahedral) arrangement

Hardware Star Tracker Sun Sensor RWAMass of System/kg 7.35 1.8 7.2Peak Power of System/W 30 0.33 33Accuracy/◦ 0.000833 1 N/A

Dimensions/mmCHU-90× 111× 139 95× 107× 35 115× 115× 86DHU-155× 210× 56

Momentum Capacity/Nms N/A N/A 1

Table 4.6.3: Key Specifications of Attitude Determination and Control Hardware


4.6.5 Hardware Application and Resultant Thruster Requirements

Slewing using the RWA for any rate > 0.05◦s−1 requires extra structural reinforcement

leading to an increase in mass. Therefore the slew rates given in Table 4.6.1 suggest

that the RWA should only be used for off-nadir data collection slews. This means that

RWA will only build up angular momentum due to these regular slew movements and

disturbance torque correction. Using a safety factor of two, the momentum storage

due to disturbances per day is calculated as,

hD = TL × 86400× 2 = 0.263Nms. (4.6.7)

The torque required to perform the off-nadir data collection slews is calculated using,

TN = 4θI

t2(4.6.8)

where θ is the angle of slew in radians (5◦ × π180

), I is the greatest moment of inertia

value after transfer (17.26 kgm2) and t is the time taken for the slew (60s). Therefore,

TN = 4× 5× π

180

17.26

602= 1.67× 10−3Nm. (4.6.9)

The momentum storage per day due to off-nadir slews is thus,

hN = 1.67× 10−3 × 10× 60 = 1.004Nms (4.6.10)

The total momentum storage per day is 1.27 Nms so in order to stay well within the

safe operating margin (80% of the wheels momentum capacity of 1Nms) momentum

dumping will occur once per orbit (4.43 times per day). External thrusters are used

to dump momentum and to perform transfer, communications and eclipse slews. The

thrusters specified for this transfer method are capable of providing a nominal thrust

of 1 N. If we assume that for each slew there is an acceleration and deceleration pulse

that lasts for 5% of the total slew time and that a momentum dumping pulse lasts for

1 s, then we can calculate required thruster forces and thus the impulse requirements

on the system. Using,

F =R

0.05× T

π

180Ix (4.6.11)

where R is slew rate, T is slew time, I is the greatest principle moment of inertia

after transfer (17.26kgm2) and x is the corresponding moment arm for this moment

of inertia value. These forces multiplied by the time over which they are applied

gives the impulse requirements of the thrusters. Force and impulse requirements for

momentum dumping and each slew type are given in Table 4.6.4.

All force values are below 1N and so within thruster limits. The outputs that are

given to ensure that the thrusters have been sized correctly are given in Table 4.6.5.


Thruster Action Required Force/N Impulse RequirementEclipse Slew 0.602 18.07Ns/yearMomentum Dumping 0.570 626.18Ns/yearTransfer Slew 0.267 9.61NsCommunications Slew 0.050 274.88Ns/year

Table 4.6.4: Impulse and Thrust Requirements for the direct chemical transfermethod.

Parameter RequirementTransfer Impulse 9.61NsYearly Impulse 919.15NsLargest Thrust 0.602NMaximum Yearly Burn Time 988.27sFrequency of Momentum Dumping 1 per orbit

Table 4.6.5: Thruster Requirements for the direct chemical transfer method

4.7 Thermal Control Subsystem

Spacecraft components can only operate within certain temperature ranges — this

will vary from component to component. Most operate around room temperature

( 293K) as this is the condition they are designed, built and tested in [62]. The

thermal subsystem of a spacecraft ensures that each component remains within its

operating range from launch through to end-of-life.

Thermal analysis at each stage of the satellite’s life will be considered in order to

determine the best thermal control subsystem. For the initial analysis there will be

assumptions made about each stage of the mission.

4.7.1 Primary Assumptions

1. Launch

• The maximum aerothermal flux from the launcher will last less than one

minute. [61]

• The Sun and Earth will provide negligible thermal inputs, as the aerothermal

flux is so high at launch.

• An average aerothermal flux will be used to calculate the temperature for

the next fifteen minutes — based on Ariane 5 data.

• The internal power dissipation will be negligible at launch.

4.7 Thermal Control Subsystem 47

2. GTO

• The Sun, Earth and internal power dissipation will be the main thermal

influences on the spacecraft, other celestial bodies will have negligible effect.

[62]

• In eclipse the only external thermal input will be the Earth’s infrared

emissions. [62]

• Before eclipse the temperature of the spacecraft will be at the maximum

equilibrium temperature. [65]

• Internal dissipation will be found by power losses due to inefficiencies of

the subsystems in operation in this orbit, assumed to be 0.4 of the total

power usage.

• Radiation experienced by the satellite in GTO will be negligible due to the

short time spent there.

3. Transfer

• Internal dissipation are found by power losses due to inefficiencies of the

subsystems in operation at transfer, assumed to be 0.4 of the total power

usage.

• There will be no eclipse periods during the transfer.

• The minimum equilibrium temperature during the transfer will be just as

the spacecraft leaves GTO.

• The maximum equilibrium temperature during the transfer will be just

before orbital insertion at the Moon.

4. Lunar Orbit

• The solar flux felt at the Moon will be the same as that at the Earth. [63]

• The Sun, Moon and internal power dissipation will be the main thermal

influences on the spacecraft; other celestial bodies will have negligible effect.

[62]

• During eclipse, the only external thermal input will be the Moon’s infrared

emissions. [62]

• Before eclipse the spacecraft will be at its equilibrium temperature. [65]

• Internal dissipation will be found by power losses due to inefficiencies of

the subsystems in operation, assumed to be 0.4 of the total power usage.


4.7.2 Method of Calculations

1. Equilibrium Temperatures

To size the thermal subsystem, the equilibrium temperatures of the spacecraft

at each stage of it’s mission will be computed. The power balance used to find

the equilibrium power of the satellite is

Qeq = Pinternal +Qexternal, (4.7.1)

where Pinternal is the internally dissipated power from the subsystems andQexternal

is the sum of the external thermal inputs, which will vary with the mission phase.

At launch, Qexternal = Qlaunch, the aerothermal heating power imposed by the

launch environment. In GTO,

Qeq = Pinternal +QearthIR +Qearthalbedo +Qsolar, (4.7.2)

where the sum of the infrared emission from the Earth, QearthIR, Earth albedo,

Qearthalbedo and the solar input, Qsolar are equal to Qexternal. In eclipse around

the Earth, the albedo and solar inputs are zero. In lunar orbit, it would then

follow that

Qeq = Pinternal +QmoonIR +Qmoonalbedo +Qsolar (4.7.3)

where QmoonIR and Qmoonalbedo are the Moons infrared emission and albedo re-

spectively. Again, in eclipse, the solar and albedo inputs are zero. The individual

terms in the equations are given by

Qlaunch = ϵqlaunchAsurface, (4.7.4)

QearthIR = ϵqearthAprojF, (4.7.5)

Qearthalbedo = αqsolarAprojρearthF, (4.7.6)

QmoonIR = ϵqmoonAprojF, (4.7.7)

Qmoonalbedo = αqsolarAprojρmoonF, (4.7.8)

where ϵ and α are the emissivities and absorptivities of the surface materials

respectively, Asurface is the surface area of the spacecraft (6 m2) and Aproj is

the projected surface area seen by the celestial body. The term F = Rorb

Rbody

2

where Rorb and Rbody are the radii of the orbit and celestial body respectively.


The constants in the equations above are given in Table 4.7.1 and are used

throughout the calculations. [62] The equilibrium temperature is then found by

Qeq = ϵσAsurfaceT4eq, (4.7.9)

where σ is the Stephan-Boltzmann constant, 5.67x108Wm−2K−4. [62]

Symbol Name Valueqlaunch Launch aerothermal flux (W/m2) 1135 [61]qsolar Solar flux (W/m2) 1350 [62]qearth Earth infrared emission (W/m2) 270 [62]ρearth Earth albedo 0.35 [62]qmoon Moon infrared emission (W/m2) 430 [63]ρmoon Moon albedo 0.1362 [64]

Table 4.7.1: Constants used in thermal balance equations

2. Launch and Eclipse Temperatures

The equilibrium temperatures calculated for the launch and eclipse environments

may not be reached, instead the satellites temperature will move towards the

Teq, with the actual temperature reached being dependent on the time duration

of the conditions.

The real temperature in the launch environment is found by using

t

τ= ln

1 + z

1− z+ 2arctan(z), (4.7.10)

where t is the duration of the launch (or any heating conditions),z is the ratio

of the actual temperature to the equilibrium temperature, TTeq

and τ is the time

constant of the satellite. The flux felt by the satellite at launch can be seen in

Figure 4.7.1.[65]

For eclipse, a similar equation can be used to find z,

t

τ= ln

z + 1

z − 1+ 2 arctan(z), (4.7.11)

and τ is easily calculated by knowing the properties of the satellite coatings,

τ =mcp

4ϵσAsurfaceT 4eq

, (4.7.12)

where m and cp are the mass and specific heat of the coating respectively. With

a numerical solver, the value of z can be found at the end of the launch and

eclipse conditions. This will be used to find T. [65]

3. Heater Sizing In order to keep the satellite at an operational temperature,

heaters may be needed to warm components during eclipse times. The power,P ,


required by the heaters is found by

P = −kA∆T, (4.7.13)

where k is the thermal conductivity of the heater. [63]

Figure 4.7.1: Aerothermal Flux Felt at Launch [61]

4.7.3 Equilibrium and Eclipse Temperatures

A thermal coating of aluminised kapton is used to keep the satellite warm. It covers

the surface of the spacecraft that is not otherwise covered with solar arrays. The

properties of these materials can be seen in Table 4.7.2. The calculated temperatures

reached for each mission stage in Table 4.7.3.

Property Aluminised Kapton Solar Arrayα 0.4 [63] 0.88 [63]ϵ 0.63 [63] 0.8 [63]Mass (kg) 1.9 2.4Area (m2) 3.97 2.03

Table 4.7.2: Properties of Surface Coatings


Mission Stage Temperature (K)Launch 302.3GTO Daylight 283GTO Eclipse 190.9Transfer 299.5Lunar Orbit Daylight 299.5Lunar Orbit Eclipse 206.2

Table 4.7.3: Temperatures Reached at Each Mission Stage

Component MinimumTemperature(K)

MaximumTemperature(K)

Propellant Tank 278 382Oxidiser Tank 266 291PCDU 258 323Baterries 288 303On Board Computer 253 323Camera 269 313

Table 4.7.4: Component Operating Temperatures

4.7.4 Thermal Control Decisions

Based on these temperatures the spacecraft will on the whole be operating within its

operational range. During the GTO and lunar eclipses however, the temperature will

fall below acceptable operational temperatures and so heaters are needed to keep the

spacecraft components warm.

The acceptable temperatures for some components can be found in Table 4.7.4, and

the kapton heaters used have a thermal conductivity of 0.16 W/m2. The heaters are

sized for the highest required heating power, i.e. the coldest conditions (in GTO). The

heaters are strategically placed so only the essential components are kept at their re-

spective operating temperatures, to minimize heater power requirement. This results

in mass and power allocations for the heaters of 0.6 kg and 20 W respectively. During

the Lunar eclipse, the temperature will not fall as low and only 5.7 W will be required.

Heaters will be controlled by thermostats and will turn on once the temperature falls

below a components acceptable temperature.

In order to ensure that the oxidiser tank stays cool during the daylight temperatures

and during launch, heat pipes will be used to transfer any heat away from the tanks

and towards the surface of the spacecraft. Heat pipes will also be required around

the propellant tank due to the temperatures reached during GTO in daylight. Fur-

thermore, to ensure that the batteries are kept warm in GTO, the heaters surrounding

them are engaged constantly.


The internal surfaces of the spacecraft that are not covered in heaters are painted

black - this is due to black paints α and ϵ values of one [63] This will maximise heat

transfer internally and help the heat be evenly distributed within the spacecraft. A

heat shield surrounding the main engine reduces the heat flow into the orbiter during

transfer and station-keeping.

4.8 Structure and Configuration

Designs for direct transfer and weak stability boundary are identical as they both

involve chemical propulsion. In fact, the internal configuration remains the same.

The amount of power required is sufficiently low enough to have body-mounted solar

panels.

The spacecraft is 3-axes stabilized because of the pointing requirements of multiple

systems : the main engine, the payload and the antenna all required to point in the

right direction.

The radiation detector, the spectrometer and the two cameras are pointing at the

Moon, while the X-band antenna needs to point towards the Earth, and is therefore

placed on the face opposite the Moon. The dust detector is placed on top of the

structure, opposite the main thruster, in order to face the direction of travel, as this

will increase the rate of particle collisions.

Attitude determination and control subsystem is the other fundamental aspect of

configuring the spacecraft.

First of all, sun sensors are disposed on each face and the star tracker system points

towards deep space.

RCS thrusters are placed three by three (one for each principal direction), on four

corners of the structure (as shown on the following cad models), all of them pointing

outwards to avoid contamination of the instruments.

Regarding thermal control, the use of body-mounted panels on 5 surfaces has

allowed all the electronics to be positioned under any surface.

The real centre of gravity of the spacecraft is actually different from the geometric

centre:

4.8 Structure and Configuration 53

Distance from geometric centre (cm)

Dx 2Dy -4Dz -10

Table 4.8.1: Gravity centre

Table 4.8.2 shows the moments of inertia of the direct chemical spacecraft at three

different points in its lifetime: at GTO with full tanks, after transfer to the Moon,

and at the end of life with empty tanks. This data is important for attitude and orbit

control. The method used to calculate the moments of inertia are identical for all

three satellites.

The position of the centre of mass of each instruments was taken, as well as the mass

of each component. The moments of inertia are calculated by using the formulas:

Ixx = mi × (Z2i + Y 2

i ) (4.8.1a)

Iyy = mi × (X2i + Z2

i ) (4.8.1b)

Izz = mi × (X2i + Y 2

i ) (4.8.1c)

The structure will be launched with the Ariane 5 vehicle. It must fit in an ejection

End of Life After transfer GTOIxx 14.58 13.95 23.34Iyy 18.32 17.26 30.58Izz 15.46 14.82 24.22

Table 4.8.2: Inertia table for Direct Chemical Method

cone with an angle of 5◦, 1200 mm height and an upper diameter of 1640 mm. Figures

4.8.1 and 4.8.2 show the spacecraft in the launch configuration mode.


Figure 4.8.1: Direct chemical propulsion model, perspective top view

Figure 4.8.2: Direct chemical propulsion model, perspective side view

4.9 Discussion of Budget Evolution and Sub-system Trade-offs 55

4.9 Discussion of Budget Evolution and Sub-system Trade-

offs

Subsystem contributions to the budgets are monitored during the design process. This

approach, in contrast to budgets that are allocated to subsystems by the systems

engineer, allows subsystem allocations to change seamlessly as the design is refined.

Negotiations of subsystem allocations only become necessary if the entire system does

not meet top level requirements such as the mission lifetime. The evolution of the

mission lifetime (see Figure 4.9.3) includes several such occasions.

The high mass of the ACS is reduced as the initially poorly defined payload re-

quirements (most importantly the tracking duration requirement, linked to the camera

exposure time) are refined. This is evident in the dry mass budget, shown in Figure

4.9.4. Similarly, the communications subsystem has attitude requirements (slew angle

and frequency necessary to point the antenna for bulk data transmissions), which,

during development, temporarily resulted in the ACS mass and power allocations

(see Figures 4.9.4 and 4.9.2) becoming so large as to cause the wet mass to exceed

the 300 kg top level requirement (see Figure 4.9.1). A consequent redesign of the

communications strategy circumvented this problem quickly. Additionally, changes in

the payload specifications with regards to the rate of data generation have influenced

the system design heavily via the communications subsystem, which has grown to

comprise 12 % and 40 % of the dry mass and power budget respectively.

Other subsystems are observed to have budget allocations that are either small

(i.e. sub 10 %), as the thermal control subsystem, or close to constant, such as the

propulsion subsystem. These subsystems show little sensitivity to changes in the re-

quirements and inputs from other subsystems. Subsystems redundancies (thermal

control: heaters, OBDH, communications: S-band, AOCS: sensors and actuators) are

included in the respective subsystem allocations. They are not found to be in conflict

with the top level requirements at any point, i.e. no redundancies had to be lost in

favour of achieving mission objectives.

Note that the power budgets in Figure 4.9.2 are for guidance only as they depict

simplified versions of the actual power requirements, which vary with the mission

phases and modes of operation (see Section 4.3). It does, for instance, not include the

propulsion subsystem power requirement during main engine or RCS thruster firings.

The operational lifetime, shown in Figure 4.9.3, is deduced from the expellant lifetime

and the lifetime of the power subsystem, which is subject to degradation. The sizing

procedures of these two subsystems are coupled to attain an equal lifetime for both,

the satellite’s operational lifetime.


0

50

100

150

200

250

300

350

Snapshots during development

Mas

s [k

g]

mdry

mexp, trsf.

mexp, AOCS

mmargin, sys.

Figure 4.9.1: Evolution of the satellite wet mass budget during development.

0

100

200

300

400

500


Pow

er r

equi

rem

ent [

W]

ppayload

pcomms.

pOBDH

pAOCS

ppropulsion

ppower

pthermal

pmargin, sys.

Figure 4.9.2: Evolution of the satellite power budget during development.

0

100

200

300

400

500

600

700

800


Life

time

[day

s]

tcommitioning

ttransfer

toperation

Figure 4.9.3: Evolution of the mission lifetime during development.

4.10 Final Mission Budget 57

0

20

40

60

80

100

120

140

160


Mas

s [k

g]

mpayload

mcomms.

mOBDH

mAOCS

mpropulsion

mpower

mthermal

mstructure

Figure 4.9.4: Evolution of the satellite dry mass budget during development.

4.10 Final Mission Budget

Subsystem Mass Component Component AllocationMargin Subtotal

[kg] (%) [kg] (%)Payload 12.1 20 14.5 10.46Communication 14.0 20 16.8 12.13OBDH 5.5 20 6.6 4.79AOCS 16.4 20 19.6 14.19Propulsion 35.3 20 42.3 30.63Power 9.8 20 11.8 8.53Thermal 3.0 20 3.6 2.60Subsystems Subtotal 115.2Structures 19.2 20 23.0 16.67Dry Mass 138.3Transfer Expellant 100.1AOCS Expellant 11.6Spacecraft Subtotal 249.9Systems Margin 20%Spacecraft Total Mass 299.9

Table 4.10.1: Final mass budget for the direct chemical transfer option.

58 5 WEAK STABILITY BOUNDARY TRANSFER METHOD

5 Weak Stability Boundary Transfer Method


5.1.1 Transfer

An STK scenario was constructed to study the ∆V requirement to perform a WSB

transfer to the moon. Two different methods have been investigated for this type of

transfer: a conventional WSB transfer via Sun-Earth L1 and another that utilizes a

lunar flyby to reach to the Sun-Earth L1 point, both illustrated in Figure 5.1.3

The two methods required very similar ∆V requirement and transfer time. They

have, however, produced a considerable difference in the final orbit: the conventional

method produced AoP of 320.0◦, compared to 4.0◦for the lunar flyby. This would

mean that the traditional WSB transfer would result in a final orbit with station-

keeping ∆V requirement similar to that of the direct chemical option, whereas for the

lunar flyby, it would require about fourth of that figure, meaning that the mission

lifetime could be extended by a factor of four. Therefore for the WSB option, the

lunar flyby method was chosen.

Approximately the same launch date was chosen as the direct chemical option. Upon

arrival, however, LEWIS is allowed to swing by. The TLE epoch is controlled to allow

control of the highest apogee altitude near L1. This allows control of the magnitude

of sun’s gravitational assist. At apogee, one targeting burn is fired to achieve the

desired PA and inclination.

5.1.2 Final Orbit

1

2

3

4

30

210

60

240

90

270

120

300

150

330

180 0

Orientation of Final Orbit for WSB Option


South - North

LEWIS Orbit

Figure 5.1.1: Final Orbit Achieved byWSB Option. AoP = 4.0◦

The final orbit achieved by WSB transfer

is near flat with AoP of 4◦. This

results in a much lower ∆V requirement

compared to the direct chemical option,

of 78m/s per year, as simulated by STK,

results shown in Figure 5.1.2.

staionkeeping The same stationkeeping

strategy as the direct chemical option is

employed: PA and AA are controlled as

soon as they drift by 10km. Since less

fuel is expended in transfer compared to

the direct chemical option, and less fuel

is required for stationkeeping per year,

LEWIS would stay on the designed orbit

for much longer, 42 months by linear extrapolation.


Similarly the power subsystem is sized to provide minimum power at this time.

Figure 5.1.2: Keeping PA at 100km ± 10km for 1year requires 78m/s. By linearextrapolation given the fuel carried, LEWIS will keep the station for 42 months inlunar orbit.

5.2 Chemical Propulsion Subsystem

The chemical propulsion subsystem for the WSB option covers the same functions as

for the direct chemical option. The two propulsion subsystems are identical in terms

of their configuration apart from the the tank dimensions. The methodology and

sizing process employed for the direct chemical option (see Section 4.2) are directly

applicable to the propulsion subsystem for WSB.

The expellant lifetime for WSB is 1252 days in contrast to 272 days for the direct

chemical option. This can be attributed to the lower expellant mass required for the

transfer and lower station-keeping delta-v in the mission orbit.

The impulse delivered by the ME during transfer is 1.13 × 105 Ns, i.e. 18.7 % of

the MEs total impulse capability of 1.67 × 106 Ns. The remainder is available for

station-keeping and RCS functions. As for direct chemical the impulse delivered by

the RCS during transfer is negligible in contrast to the 1.80 × 105 Ns capability of

a single RCS thruster. Furthermore, with the annual RCS impulse requirement of

920 Ns, neither the lifetime of the ME nor the RCS thrusters impose constraints on

the mission lifetime.

A breakdown of the dry propulsion subsystem in terms of mass is given in Table 5.2.1.


Figure 5.1.3: STK simulaitons on WSB transfers. Conventional WSB transfer viaL1(Top): Transfer ∆V = 1177m/s, Transfer Time = 101 days. And transfer via lunarflyby (Bottom): Transfer ∆V = 1048m/s, Transfer Time = 80 days. For advantagesin final orbit, the latter method is employed.


Note the miniscule differences compared to the similar dual-mode system employed

for the direct chemical transfer (see Table 4.2.1 in Section 4.2).

Component Mass Allocation[kg] (%)

Oxidiser tank 9.71 27.6Propellant tank 10.04 28.5Pressurant tank 2.34 6.6RCS thrusters (12) 3.48 9.9ME 6.03 17.1Plumbing 3.16 9.0Pressurant 0.43 1.2Propulsion system dry 35.18 100.0(incl. pressurant)

Table 5.2.1: Mass breakdown of the dry dual-mode propulsion subsystem for the WSBtransfer option.

5.3 Power Subsystem

5.3.1 Requirements

To estimate the power requirement for the spacecraft, the same approach used for

the direct chemical option, described in Section 4.3, has been used. The power re-

quirement of each instrument has been has been given, and a power plan over one

orbit has been established. However, since the eclipse happens at a different section

of the orbit, the power plan looks slightly different, as shown in Figure 5.3.1.

The average power 114.7 W; and with 10% contingency, subsystem power re-

quirement (PSR) is given to be 126.2 W. To size the power subsystem, it is assumed

that the spacecraft steadily consumes this amount of power.

Figure 5.3.1: Power profile for WSB transfer option.


5.3.2 PCDU

The magnitudes of power required and produced by this spacecraft are very similar to

that of the direct chemical option. Hence, the Small Satellite Power System by SSTL

is also used for this satellite.


It has been shown that PSR = 126.2 W. The longest possible eclipse time has also

been found to be 3620 seconds. Taking depth of discharge of 80%, The minimum

energy capacity of the battery is given by Equation (4.3.1) to be 181.7 Wh. This is

larger than the energy capacity of the VES-180 cell shown in Figure 4.3.3. Hence,

two of this battery are carried in the spacecraft, giving Ebattery = 360 Wh.


The end-of-life power for the spacecraft is given by Equation (4.3.2) to be 167.8 W.

Since the same length of time is spent for the commissioning phase in GTO, the same

amount of damage is expected during this phase as with the direct chemical option. It

is also assumed that the radiation environment in the trans-lunar space and near L1

is similar to that at the lunar orbit, the main constituent of which being the protons

from the sun and galactic cosmic rays.

Using these figures, the beginning-of-life power has been found to be 193.7 W.

The same solar cell from emcore is used for its high efficiency. Using Equation

(4.3.4), the sun-projection area of the array is found to be 0.49 m2.


Since the sun-projection area is less than 1 m2, body-mounted configuration is chosen.

As with direct chemical method, the arrays are mounted on five faces.

5.3.6 Power Profile

Using the PCDU, batteries and solar arrays, a power profile is simulated over one

orbit, results shown in Figure 5.3.1.

5.4 OBDH

Given the transfer time of about 80 days and the 10 mm thick shielding, the amount

of radiation to be received by the spacecraft is estimated to be about 5.617 krad.



Table 5.5.1 shows the final set of data rates to be used in the case of the weak stability

boundary transfer method. These values were calculated using the CDE spreadsheets

and are considered to be the final output for the communications segment.



36.45 kbps

Table 5.5.1: Final key values for the weak stability boundary option.

5.6 ADCS

5.6.1 Control Modes and Selection of Attitude Control Modes

The control modes for the WSB transfer method will be exactly the same as the

direct chemical transfer method. It will also have the same ADCS requirements for

its slewing modes and camera requirements as given in Tables 4.6.1 and 4.6.2. As a

result the argument for the attitude control method will also be the same as for the

direct chemical transfer for the WSB transfer. The conclusion of this was to use a

zero momentum three wheel reaction wheel assembly.

5.6.2 Quantifying the Disturbance Environment

The final lunar orbit for this transfer method will be very similar to the direct chemical

option. The largest surface has remained 0.5m2. Only the inertia matrix has changed

due to some changes in hardware and configuration.

This means that Ts is still 6.83× 10−7 Nm. The new disturbance torque to gravity

gradient effects can be calculated using Equation 4.6.1 by using the new K value

of 2.80kgm2 and the method outlined in Section 4.6.3. This yields a Tg value of

7.11× 10−7 Nm and thus TL = 1.39× 10−6.


As the ADCS requirements for the WSB method are extremely similar to the direct

chemical method, the same hardware has been selected with the properties as listed

in Section 4.6.4.


5.6.4 Hardware Application and Resultant Thruster Requirements

As the requirements on the ACDS are the same apart from the changes in the inertia,

the method and equations listed in Section 4.6.5 can be used as before. The total

momentum storage per day is 1.31 Nms, compared to 1.27Nms for the direct chemical

method. In order to stay within the RWA’s operating limits, momentum dumping

will occur similarly once per orbit (4.43 times per day). This yields the impulse and

thruster requirements given in Tables 5.6.1 and 5.6.2.

Thruster Action Required Force/N Impulse RequirementEclipse Slew 0.644 19.33Ns/yearMomentum Dumping 0.590 649.62Ns/yearTransfer Slew 0.467 12.61 NsCommunications Slew 0.054 294.00 Ns/year

Table 5.6.1: Impulse and Thrust Requirements for the WSB transfer method.

Parameter RequirementTransfer Impulse 19.33NsYearly Impulse 962.95NsLargest Thrust 0.644NMaximum Yearly Burn Time 988.27sFrequency of Momentum Dumping 1 per orbit

Table 5.6.2: Thruster Requirements for the WSB transfer method




The same assumptions made in Section 4.7.1 will be employed here. The transfer

method will use a lunar fly by. An additional assumption here is that the spacecraft

will not experience any eclipses during this transfer other than the short fly-by.


The same method outlined in Section 4.7.2 will be used to find the equilibrium and

transient temperatures, and to size any heaters needed.


As with the direct chemical option, aluminised kapton will be used on spacecraft

surfaces not coverd by solar arrays. The properties of the kapton and the arrays can

be seen in Table4.7.2. The calculated temperatures reached for each mission stage in

Table 5.7.1.

Mission Stage Temperature (K)Launch 302.8GTO Daylight 276.7GTO Eclipse 191.4Transfer 205.4Lunar Orbit Daylight 294Lunar Orbit Eclipse 205.4



The temperatures reached on the whole are lower than those reached for the direct

chemical option, however still mainly within operational ranges. This will require the

heaters to be operating with a higher power during GTO and lunar eclipses, as well

as during GTO daylight for both the propellant tank and batteries (see Table 4.7.4

for operating temperatures).

The spacecraft will require 0.62 kg of kapton heaters to keep the components warm

during the GTO eclipse (as this is the coldest scenario the spacecraft will face) and

this will require 23.4 W of power.


Although the equilibrium temperatures are lower for this option, the oxidiser tank

will still require heat pipes to transfer excess heat away from the tanks and towards

the surface of the spacecraft.

The internal surfaces of the spacecraft that are not covered in heaters will be painted

black - this is due to black paints α and ϵ values of one. Again, body mounted solar

arrays meant that the surface coating of the spacecraft could not be chosen entirely

for thermal control.


The configuration considerations for the WSB option are similar to those for the

direct chemical option found in section 4.8. The transfer method is achieved by using

chemical propulsion as well. Therefore, the only differences are the diameters of the

spherical tanks which only change by a few millimeters.

Direct chemical Weak stability boundary

Propellant tank 452 mm 451 mmOxidiser tank 423 mm 419 mmPressurant tank 290 mm 288 mm

Table 5.8.1: Tank diameters

The spacecraft is 3-axes stabilized because of the pointing requirements of multiple

systems, such as the main engine, the payload and the X-band antenna.

The actual centre of gravity of the spacecraft is not precisely at the geometric centre:


Dx 2Dy -4Dz -10


The moments of inertia are then almost identical to the ones for direct chemical

and given in Table 5.8.3.


End of Life After transfer GTO

Ixx 15.69 15.66 24.71Iyy 18.47 18.46 30.10Izz 16.01 15.97 24.58

Table 5.8.3: Inertia table

Figures 5.8.1 and 5.8.2 present the spacecraft used for the weak stability boundary

method, in top view perspective, and side view perspective. Each component are to

scale, and have the appropriate dimensions in the model. Some of the parts such

as structural elements, data cables, propellant and heat pipes are not shown in the

diagrams.

The structure is also launched using Ariane 5. It has to fit in an ejection cone with an

angle of 5◦, 1200 mm height and an upper diameter of 1640mm. The following figure

shows the spacecraft in the launch configuration mode.

Figure 5.8.1: WSB model perspective top view


Figure 5.8.2: WSB model, perspective side view

5.9 Discussion of the Budget Evolution due to Sub-system

Trade-offs

During the concurrent design process, the mass budget, power budget, and mission

profile all varied as subsystems were traded off against each other.

It should be noted that early in the design process, limited use was made of the

weak stability boundary CDE spreadsheet, since it was assumed that most spacecraft

properties would match those of the spacecraft using the direct chemical transfer

method. This is why in all graphs there is little variation until approximately two

thirds of the way through.

Figure 5.9.3 shows the sudden change in subsystem power requirements as the

weak stability boundary CDE spreadsheets become populated and assumed rough

values are removed. The most significant changes are in the power required by the

thermal control system, communications system, and payload.

As the thermal and communications power requirements become much larger the

5.9 Discussion of the Budget Evolution due to Sub-system Trade-offs 69

payload power decreases. Overall, the total power does not change by more than

40%. This may be due to the choice for a body mounted solar array, which inherently

limits power due to the limited array area. It can also be seen that these values vary

less between the last few iterations as the power becomes better optimised.

Figure 5.9.4 shows very little variation. The weak stability boundary transfer

method is limited in its flexibility since the spacecraft must follow such a specific

route. This means that although later in the design the time spent during transfer

has been reduced, the variation in transfer time is small.

0

20

40

60

80

100

120


Mas

s [k

g]

mpayload

mcomms.

mOBDH

mAOCS

mpropulsion

mpower

mthermal

mstructure

Figure 5.9.1: Evolution of the satellite dry mass budget during development

0

50

100

150

200

250

300


Mas

s [k

g]

mdry

mexp, trsf.

mexp, AOCS

mmargin, sys.



0

100

200

300

400


Pow

er r

equi

rem

ent [

W]

ppayload

pcomms.

pOBDH

pAOCS

ppropulsion

ppower

pthermal

pmargin, sys.


0

100

200

300

400


Life

time

[day

s]

tcommitioning

ttransfer

toperation




[kg] (%) [kg] (%)Payload 12.1 20 14.5 10.31Communication 14.0 20 16.8 11.97OBDH 5.5 20 6.6 4.72AOCS 16.4 20 19.6 13.99Propulsion 35.2 20 42.2 30.11Power 11.3 20 13.6 9.70Thermal 3.0 20 3.6 2.54Subsystems Subtotal 116.8Structures 19.5 20 23.4 16.67Dry Mass 140.2Transfer Expellant 96.6AOCS Expellant 13.2Spacecraft Subtotal 250.0Systems Margin 20 %Spacecraft Total Mass 300.0

Table 5.10.1: Final mass budget for the chemical WSB transfer option.

71

6 Low Thrust Transfer Method


6.1.1 Transfer

-5

0

5

x 105

-6

-4

-2

0

2

4

x 105

-1.5

-1

-0.5

0

0.5

1

1.5

x 105

X-distance from Earth Centre [km]

Transfer Trajectory of LEWIS in Low Thrust Option

Y-distance from Earth Centre [km]

Z-distance from Earth Centre [km]

Van Allen Belt

Figure 6.1.1: Low Thrust Trajectory

To predict the ∆V requirement

using an ionic engine, the EMT

software provided by Dr. Hugh

Lewis was used. Input pa-

rameters used are outlined in

Table 6.1.1, together with the

resultant fuel-mass used and the

transfer time. The resultant

trajectory is then shown in

Figure 6.1.1.

According to the results produced

by the EMT software, it takes

about 300 days since the ac-

tivation of burn for LEWIS to achieve a semimajor axis higher than the outer

boundaries of the Van Allen belts. As shown in Section 6.3, the solar arrays for this

option have been scaled from the SMART-1 mission. This has allowed the scaling of

their degradation from SMART-1 as well. [53]

Table 6.1.1: Inputs and Outputs of EMT software

Initial Conditions Final Conditions

Semimajor Axis [km] 24661.14 Semimajor Axis [km] 3746.75Eccentricity 0.716228 Eccentricity 0.51131

Inclination [deg] 7 Inclination [deg] 90Ascending Node [deg] 0 Ascending Node [deg] 0

Argument of Perigee [deg] 178 Argument of Perigee [deg] 90True Anomaly [deg] 0 True Anomaly [deg] 0

Depart Date 00:00:0023/01/2013

Arrival Date 15:50:2423/01/2015

Transfer Fuel [kg] 32.7 Transfer Time [days] 730.7

72 6 LOW THRUST TRANSFER METHOD

6.1.2 Final Orbit

1

2

3

4

30

210

60

240

90

270

120

300

150

330

180 0

Orientation of Final Orbit for Low Thrust Option


South - North

LEWIS Orbit

Figure 6.1.2: Final Orbit Achieved by lowthrust option. AoP = 0◦

Low thrust transfer method has allowed

free choice of the final orbit. The desired

orbit with AoP of zero is hence achieved.

Stationkeeping As AoP of zero is

achieved, PA and AA will only slowly

drift. The first stationkeeping maneuvre

is required at the 17th week since TLI,

as shown in Figure 6.1.3.

The stationkeeping strategy for the

low thrust option differs to those

employed by the chemical options, due

to the difference in the thrust levels

achievable by the engines. For this option, only PA is controlled, and AA is allowed

to drift freely, as the stationkeeping burn would take about an hour, and it is not

desirable to use this time to control AA, which could otherwise be used to perform

observations near the periselene. Burns are performed near aposelene to keep the

periselene well above surface, as there is plenty of time when the spacecraft is orbiting

slowly.

The tolerance PA is allowed to drift is calculated by the following steps:

1. On 100 km × 3600 km orbit, find time taken for True Anomaly to proceed from

170◦to 190◦. This is the time used to perform the aposelene maneuvre.

2. Use first guess for PA tolerance (10 km was used), to calculate ∆V required at

aposelene

3. Use rocket equation to calculate mass of propellant required

4. Use mass-flow rate of the engine used to calculate the time taken for the burn,

compare with value obtained in 1.

5. Repeat 2-4 to find allowable tolerance for PA.

As a result, tolerance of 7.5 km was found. STK simulation was then performed using

impulsive maneuvres at aposelene to find how many burns are required per year, the

result shown in Figure 6.1.3. Annual ∆V requirement for stationkeeping was found

to be 8 m/s; much lower than the values for the chemical options. The difference is

due to the difference in the stationkeeping strategies, and the relative stability of the

orbit obtained by the low thrust option. By linear extrapolation with the fuel carried

on board, it is estimated that the spacecraft will stay in designed orbit for about 14

6.2 Electric Propulsion Subsystem 73

years. It may well be that other subsystem will fail before this point, however, it took

only 1 kg of fuel to achieve the 14 years, and the spacecraft weighs 280 kg in total,

allowing for additional payload or contingencies.

Figure 6.1.3: Keeping PA at 100 km ± 7.5 km for 1year requires 8.1 m/s. By linearextrapolation given the fuel carried, LEWIS will keep the station for 13.9 years inlunar orbit.

6.2 Electric Propulsion Subsystem

The choice for the LEWIS mission is the QinetiQ-manufactured Kaufman type T5

thruster shown in Figure 6.2.1. The ion engine is has grids 10 cm in diameter. Direct

current is discharged between the hollow cathode and cylindrical anode to ionise the

Xenon. [54] A simple side view of the gridded ion thruster is shown in Figure 6.2.1.

The ion engine assembly comprises four major components, configured in the ar-

chitecture shown in Figure 6.2.2, resulting in a 30.7kg dry mass.

The gridded ion thruster T5, has the key parameters shown in Table 6.2.1.

The Proportional Xenon Feed Assembly (PXFA), designed by MOOGBradford,

regulates and maintains the flow of Xenon from its tanks to the main cathode and

neutraliser to feed the T5 thruster.[56] The key specifications the PXFA are shown in

Table 6.2.2.

The Ion Propulsion Control Unit (IPCU) designed by Astrium provides the

required voltage and current to the thrusters, as well as measuring the temperature,

and controlling the propellant flow.[57] The key specifications for the IPCU are given

in Table 6.2.3.

Xenon Propellant and propellant tank for T5 thruster - xenon has been

chosen for propellant, due to its high performance, non-toxicity and good storage


Figure 6.2.1: QineticQ manufactured Kaufman type T5 Thruster

Figure 6.2.2: Electrical Propulsion Architecture

Parameter ValueMass 2.95kg (including adjustable mounting bracket)Thrust 1-20 mN (±12µN)Power 55 to 585 WSpecific Impulse 500 to 3500 sTypical efficiency 66% (approx.)Exhaust velocity 31.74kms−1

Dimensions 180× 200× 100.dia

Table 6.2.1: Key parameters of the gridded ion thruster.[55]

Parameter ValueMass 7.5 kgInput pressure range 5:25 barRegulated Pressure Range 2.5 barMain Mass flow rate 0.63mgs−1

Dimensions 150x 200 mm x 350 mm

Table 6.2.2: Key parameters of the Proportional Xenon Feed Assembly (PXFA).[55]

6.2 Electric Propulsion Subsystem 75

Parameter ValueMass 17.5kgInput Voltages 22- 37VMaximum Input Current 37A @ 22 VDimensions 380× 270× 205mm

Table 6.2.3: Key parameters of the Ion Propulsion Control Unit (IPCU).[55]

properties. It is kept under 150 bar in a supercritical phase, above the critical point

of 298 K in a propellant tank made of metal composite, as illustrated in Figure

6.2.3. In order to keep the Xenon in that phase and avoid fluid condensation or liq-

uefaction, the propellant tank is covered with a thermal blanket. It is placed along

the central orbiter axis inside an aluminium cylinder, insulated from the rest of the

orbiter systems.[58, 59, 60]

The Xenon required for the 730.65 days transfer and 6 months mission orbit was

Figure 6.2.3: The Xenon Phase Diagram showing the liquid, gaseous and supercriticalphases

estimated from EMT software to be 34.29 kg. This includes a 2% margin due to the

residual propellant left in the tank.[59] The Xenon tank mass of 2.76 kg is calculated

by linearly scaling from the SMART-1 mission, which has a propellant tank mass of

7.7 kg and tank volume of 50 l. This gives a tank density of 0.154 kg/l. A margin

of 10% for the propellant tank volume : 17.98 l : is included in the calculation, ac-

counting for the ullage of the gas, or the gas left in the tank during the fuelling.


6.2.1 Low Thrust Reaction Control System

The low thrust option employs a conventional mono-propellant RCS, comprised of 12

hydrazine thrusters (see Table 4.2.2 (a)). The hydrazine is stored in a spherical tank

with diaphragm for propellant capture. The tank mass is estimated to be the mass

of a spherical Titanium tank of constant wall thickness and a 50 % addition for the

diaphragm. The propellant is blowdown-pressurised keeping the system simple and

reliable.

Of the 7.13 kg of stored hydrazine, 1 kg is allocated for station-keeping by repurposing

the RCS thrusters (if necessary), the remainder is for RCS functions. After the transfer

(RCS impulse: 1910 Ns) a RCS impulse of 11332 Ns is left, i.e. a RCS propellant

lifetime of 4 years in the mission orbit (the annual RCS impulse requirement is 2730

Ns).

6.3 Power Subsystem

6.3.1 Requirements

To estimate the power requirement, similar method has been used as with the direct

chemical and WSB options: instruments have been identified and a power plan has

been constructed, as shown in Figure 6.3.1. The average power requirement over an

average orbit is found to be 174.6 W, and with 10% contingency, the subsystem power

requirement is (PSR) is found to be 210.0W. This value is used to size the batteries.

However, this figure is less than the power required by the propulsion subsystem alone,

which is 702.0 W.

Hence to size the power subsystem, the power requirement at the cruise phase is

found, which includes the peak powers of propulsion subsystem, OBDH and AOCS

subsystem, which sum up to 826.9 W. This value is used to size the solar arrays.

Figure 6.3.1: Power profile for low thrust transfer option.


6.3.2 PCDU

As shown below, The maximum power produced by the solar arrays is still less than

1.6 kW. For this reason, Small Satellite Power System by SSTL is used for this satellite

as well.


From the power plan, PSR has been found to be 210.0 W. The batteries’ requirement

is to provide this power during the eclipses. The longest eclipse for the given orbit

is found to be 3627 seconds. Taking the VES-180 cell with DoD of 80%, Equation

(4.3.1) is used to find the minimum energy capacity of the battery (Emin) to be 281.3

Wh. Since VES-180 features 180 Wh, two of this battery is are placed on board.


Due to the high cruising power requirement, the solar arrays are going to be deployed.

To be able to accurately estimate the mass of the deployable structure, the solar array

for this spacecraft is scaled from spacecraft data made available from the SMART-1

mission.

It has been found that cruising power requirement is greater than the subsystem

power requirement, therefore, the solar arrays are sized to the cruising power re-

quirement of 826.9 W. According to the solution provided by the EMT software,

the low thrust transfer takes 730.7 days. The solar arrays must then provide the

cruising power requirement for this duration. The solar arrays on-board the SMART-

1 spacecraft has degraded to 87.4% of the BOL power in its transfer time of 532 days.

Scaled from this value, the solar arrays on board the spacecraft degrades by 83.1%.

[53]

Taking also the lunar radiation environment, it has therefore be found that the

beginning-of-life power is 994.9 W.

Using Equation (4.3.4), the sun-projection area of the solar aray is then found to be

2.5 m2.


It has been found that the sun-projection area required is larger than the projection

area available on the spacecraft surface. Therefore, a deployable solar array is used.

To place the CG within the spacecraft body, two wings are mounted on the opposite

sides. The arrays are rotatable with an actuator, which receives commands from the

OBDH subsystem. [23]


6.3.6 Power Profile

Using the components specified, the power profile has been compiled, shown in Figure

6.3.1.

6.4 OBDH

During the entire lifetime, by considering the 7 mm thick shielding and the transfer

time of 730.7 days, the amount of radiation received by the OBDH is estimated to be

about 7.759 krad [26]. Hence the OBDH subsystem, which is certified to stand up to

100 krad, is considered to be latch-up immune.

Aside from the general OBDH configuration, the spacecraft would need a solar-

arrays pointing-mechanism controller. Being this an analog device, it can be interfaced

with the OBC and driven by an apposite software. The code will take as input the

information from Sun sensors and, knowing the relative position of the spacecraft

with respect to the Sun, it will move the arrays in order to illuminate the maximum

available surface.

Given the enhanced spacecraft power availability, up to 17 pictures per camera

per orbit can be downlinked to the ground station. This means that the amount of

data collected and compressed during each orbit rises from 110.27 MB/orbit to 169

MB/orbit, which is still a data size storable and manageable by the OBC.


Table 6.5.1 shows the final set of data rates to be used in the case of the low thrust

transfer method. These values were calculated using the CDE spreadsheets and are

considered to be the final output for the communications segment.



36.45 kbps

Table 6.5.1: Final key values for the low thrust option.

6.6 ADCS 79

6.6 ADCS

6.6.1 Control Modes and Selection of Attitude Control Mode

The control modes for the low thrust method are the same as for direct chemical

apart from the transfer mode. Instead of only needing to align for three burns during

transfer, the orbiter’s thruster must constantly be aligned to the desired thrust axis

during the low thrust transfer method. The remaining slew and camera requirements

given in Tables 4.6.1 and 4.6.2 remain the same. This implies that the attitude control

method for the low thrust transfer method will be the same as for the direct chemical

and WSB transfer. The conclusion of this was to use a zero momentum three wheel

reaction wheel assembly.

6.6.2 Quantify the Disturbance Environment

The total disturbance torque when in final orbit can be calculated using the method

outlined in Section 4.6.3. The use of solar arrays has meant that the largest surface

array is now 6.3 m2 and K is now 21.65 kgm2. This gives Ts = 4.31× 10−6 Nm

and Tg = 5.50× 10−6 Nm, thus yielding a total lunar disturbance torque of TL =

9.80× 10−6 Nm.

During the transfer phase, disturbance torques caused by magnetic and gravity gradient

effects from the Earth must also be considered. The magnetic effects can be quantified

using

Tm =2DM

R3(6.6.1)

where D is the residual dipole of the orbiter in (which we will assume is 1 Am2), R is

the orbit Radius and M is the magnetic moment of the Earth (7.96× 1015 tesla.m3).

The torque due to the gravity gradient effects of the Earth can be calculated using

Equation 4.6.1, where µ = 3.99× 1014 Nm2/kg and θ = 45◦.

As the orbiter travels towards the moon, the effects of the Earth will diminish as

the effects of the Moon increase. The worst case scenario would be for the orbiter

and the Moon to remain aligned in their orbit around the Earth. By observing co-

ordinate data produced during calculations as described in Section 6.2, it is estimated

that the transfer time will be 730.66 days taking 508 orbits of the Earth. Assuming

each orbit is circular, the radii of the orbits must increase by an average of 0.48%

each orbit, to ensure the orbiter reaches its final lunar orbit. This implies that the

Earth caused disturbance torques will decrease and disturbance torques due to the

Moon will increase, by 1.44% with every increasing circular orbit. If it is assumed

that the solar radiation pressure torque remains constant at Ts = 4.31× 10−6, then

the variation of the total and separate disturbance torques over the transfer time can

be plotted as in Figure 6.6.1. From these results we can infer that the average total


Figure 6.6.1: Disturbance Torque variation during transfer

disturbance torque during transfer is TT = 2.52× 10−6 Nm.


As the ADCS requirements for the low thrust method are similar to the other two.

Initially the same hardware was selected as described in Section 4.6.4, however, after

preliminary calculations it was clear that the angular momentum storage of the RWA

would not be sufficient. As an alternative the Smallwheel 200SP reaction wheel,

designed by SSTL was chosen. The key specifications for this wheel are given in

Table 6.6.1.

Parameter RequirementMass of System/kg 20.8Peak Power of System/W 207.93Dimensions/mm 240.dia×90Momentum Capacity/Nms 12

[htbp]

Table 6.6.1: Key Specifications for the SSTL built Smallwheel 200SP

6.6.4 Hardware Application and Resultant Hardware Requirements

TT is multiplied with the duration of transfer (63129023.65s) and a safety factor of two

(to take into account disturbances caused by the Van Allen belts and other difficult


to predict sources). This allows us to get the total stored angular momentum during

transfer, hT = 159.09 Nms. The total momentum storage per day is therefore 0.22

Nms so in order to stay within safe operating margins momentum dumping would

have to occur at least once every 40 days. However, after experimental calculations it

is clear that the nominal thruster force of 1 N will be the limiting factor. To remain

within thruster operating limits it is necessary to perform a 1 s momentum dump

pulse for each axis every two days.

The total momentum storage per day during operation is 4.14 Nms. In order to

stay within the RWAs operating limits momentum dumping would need to occur every

two days. However as thrusters can produce a maximum thrust level of 1N, momentum

dumping must occur once per orbit. Now using the method and equations listed in

Section 4.6.5, the requirements on the ADCS and the thruster requirements can be

determined. These are given in Tables 6.6.2 and 6.6.3

Thruster Action Required Force/N Impulse RequirementEclipse Slew 0.655 29.47Ns/yearOperational Momentum Dumping 0.926 2028.70Ns/yearTransfer Momentum Dumping 0.871 1909.02 NsCommunications Slew 0.123 672.23 Ns/year

Table 6.6.2: Impulse and Thrust Requirements for the low thrust transfer method.

Parameter RequirementTransfer Impulse 1909.02NsYearly Impulse 2730NsLargest Thrust 0.926NMaximum Yearly Thruster Burn Time 1368sTransfer Thruster Burn Time 1096sFrequency of Transfer Momentum Dumping 0.5 per dayFrequency of Operational Momentum Dumping 1 per orbit

Table 6.6.3: Thruster Requirements for the low thrust transfer method



The assumptions made in Section 4.7.1 will also apply to the low thrust case. The

transfer method will differ however. Eclipses will be experienced during the transfer

and the assumption is that the longest of these will last two hours, where the spacecraft

will fall to its lowest temperature [66].



The method given in 4.7.2 will be used to find the equilibrium and transient tem-

peratures, and to size any heaters needed.


As the solar arrays are not body mounted for this satellite, the surface properties have

changed. This satellite will radiate far more heat than the other two options due to

the high power usage of the propulsion system.

For the spacecraft to reach acceptable equilibrium temperatures for its mission,

Multi Layer Insulation (MLI) and white paint (with a high ϵ) have been chosen as

coatings. The proportion of MLI-to-paint was determined on a trial and error basis,

and evolved as the design of the satellite changed. The properties of the MLI and

white paint are compared in Table6.7.1. The calculated temperatures for each mission

stage are given in Table 6.7.2.

Property MLI White Paintα 0.14 [67] 0.26 [63]ϵ 0.035 [67] 0.83 [63]Mass (kg) 1.92 0.23Area (m2) 4 2

Table 6.7.1: Properties of Surface Coatings

Mission Stage Temperature (K)Launch 301.3GTO Daylight 239.9GTO Eclipse 212.4Transfer 151.3Lunar Orbit Daylight 298.3Lunar Orbit Eclipse 273.5



The low thrust spacecraft has the highest equilibrium temperatures of all the spacecraft,

and it does not go above the operating temperatures for the components (see Table

6.7.3. Heaters however will still be required for some stages of the mission. The

main use of heaters will be during the transfer to the moon where the spacecraft will

drop to its lowest temperature. 0.62 kg of heaters consuming 23.4 W of power will


be required. Once in lunar orbit however, only 0.6 W will be needed for the lunar

eclipses. The propellant tank is coated in its own thermal blanket, as discussed in

Section 6.2.

Component MinimumTemperature(K)

MaximumTemperature(K)

Propellant Tank 233 433Hydrazine 275 386PCDU 258 323Baterries 288 303On Board Computer 253 323Camera 269 313

Table 6.7.3: Component Operating Temperatures for Low Thrust Option

Again, to promote uniform temperatures the interior of the satellite will be painted

black. This option has allowed the surface coating of the entire satellite to be chosen

for thermal reasons, which lead to minimal active thermal control in the lunar orbit

and allows for the available power to be allocated to other subsystems.


For all three spacecrafts, an almost identical main structure is used in order to increase

the reusability of the design.

The designs for chemical and electrical propulsion only differ in terms of type of

fuel tanks used and solar panel area.This mission requires more power and body-

mounted panels would not be sufficient, therefore deployed solar arrays are used

instead. Moments of inertia are therefore affected:

End of Life After transfer GTO

Ixx 37..21 37.65 38.58Iyy 15.56 16.65 19.08Izz 32.80 34.25 34.95

Table 6.8.1: Inertia table

The xenon tank, the ion propulsion control unit and the proportional xenon feed

assembly replace the oxidiser, and fuel tanks present in the chemical designs. A small

propellant tank (diam : 256mm) is added for the RCS thrusters.


The radiation detector, the spectrometer and the two cameras are pointing at the

Moon, while the X-band antenna needs to point towards the Earth, and is therefore

placed on the face opposite the Moon. The dust detector is placed on top of the

structure, opposite the main thruster, in order to face the direction of travel, as this

will increase the rate of particle collisions.

The two deployed solar arrays are positioned on the two remaining faces. They are

rotatable with an actuator, which receives commands from the OBDH subsystem.

The centre of gravity is being kept quite close to the geometric centre of the spacecraft.


Dx 0.8Dy 5Dz -5


Figures 6.8.1 and 6.8.2 present the spacecraft used for the low thrust transfer, in

top view perspective, and side view perspective. Each component are to scale, and

have the appropriate dimensions in the model. Some of the parts such as structural

elements, data cables, propellant and heat pipes are not shown in the diagrams.

As for the other two satellites, the structure will be launched with the Ariane 5

vehicle. It must fit in an ejection cone with an angle of 5◦, 1200 mm height and an

upper diameter of 1640mm. This represents another challenge in terms of fitting the

spacecraft into the launcher. The following figure shows the spacecraft in the launch

configuration mode.


Figure 6.8.1: Low thrust propulsion model, perspective top view

Figure 6.8.2: Low thrust model, perspective side view


6.9 Discussion of the Budget Evolution due to Sub-system

Trade-offs

Figures 6.9.1, 6.9.2, 6.9.3 and 6.9.4 show the evolution of the budgets taken from

the spreadsheets. As shown in these figures, the trade-off made by the propulsion

subsystem has had the biggest impact on these budgets. The trade-offs are outlined

below.

After replacing the PPS-1350-G ion engine with T5, the power dropped almost

twice down to 1300W, but the mission life time increased with almost 497.4 days.

This decision was made to accommodate the heavier reaction wheels needed to fulfil

the pointing requirements of the mission, as discussed in Section 6.6.3.

Key parameters with some of the researched electrical engines are presented in Table

6.9.1

PPS-1350-G NSTAR T5 RITA150 RIT-22Thrust (mN) 20 92.7 20 150 250Power (W) 1500 2522 585 4300 5000

Isp (s) 1660 3127 3500 5000 4500Mass (kg) 5.3 8.33 2.95 6 7

Table 6.9.1: Comparison of ion engines [70][71][72][73][74][75][76][77][78][79][80]

Simulations in EMT software were made in order to find the fastest transfer

time. A PPS-1350-G and T5 were simulated; the rest of the engines have not been

considered, due to their high power consumptions. As a first iteration, a PPS-1350-G

thruster was chosen for its shorter transfer time of 223.3 days, compared to, 730.7 days

for T5. Later on, it was discovered that a heavier set of reaction wheels were required

to satisfy the pointing requirements. Such wheels could not be accommodated due to

the mass restrictions. Therefore, the PPS-1350-G thruster was replaced with T5 ion

engine configuration. Ultimately this lead to a spacecraft weighing 282kg.

The EMT software over-estimates the fuel mass required, due to the assumption of

constant thrust during transfer. In reality, the engine will be fired only half-the orbit

spiralling out/in. Therefore, it was assumed that the 34.29 kg of Xenon, including

1kg for station keeping, will be sufficient to fulfill the mission.

Having such a long transfer time increases the risk of S/C systems damages due

to radiation exposure. The damage on the subsystems and solar arrays have been

accounted for in their respective sections.

6.9 Discussion of the Budget Evolution due to Sub-system Trade-offs 87

0

50

100

150

200

250


Mas

s [k

g]

mpayload

mcomms.

mOBDH

mAOCS

mpropulsion

mpower

mthermal

mstructure

Figure 6.9.1: Evolution of the satellite dry mass budget during development

0

50

100

150

200

250

300


Mas

s [k

g]

mdry

mexp, trsf.

mexp, AOCS

mmargin, sys.



0

500

1000

1500

2000

2500

3000


Pow

er r

equi

rem

ent [

W]

ppayload

pcomms.

pOBDH

pAOCS

ppropulsion

ppower

pthermal

pmargin, sys.


0

100

200

300

400

500

600

700

800

900

1000


Life

time

[day

s]

tcommitioning

ttransfer

toperation


.

6.10 Final Mission Budget 89

A trade-off whether to liquify Xenon when stored was made. Liquified storage

reduces the overall S/C mass approximately by 1%, however, a cryogenic system would

require complex thermal control, which adds to the risk of failure. This surpasses the

advantage of having a slightly lighter S/C. [69] Hence, it was decided to keep the

Xenon in super critical condition, as explained in Section 6.2.

Different pressure for the Xenon tanks were considered, 150 bar and 170 bar.

Higher pressure would reduce the volume of the Xenon tank: 17.98 l at 150 bar and

17.56 l at 170 bar for the T5 thruster. However, higher pressure requires thicker tank

walls. Overall, this would increase the mass of the subsystem. Therefore, a 150 bar

Xenon tank was chosen.

Given fuel mass, the tank volume of 17.98 l was calculated with the help of NIST

Chemistry WebBook, National Institute of Standards and Technology.[68]



[kg] (%) [kg] (%)Payload 12.1 20 14.5 7.47Communication 14.0 20 16.8 8.67OBDH 5.5 20 6.6 3.42AOCS 35.6 20 42.8 22.11Propulsion 30.7 20 36.9 19.06Power 33.3 20 39.9 20.63Thermal 3.2 20 3.8 1.97Subsystems Subtotal 161.2Structures 26.9 20 32.2 16.67Dry Mass 193.5Transfer Expellant 33.3AOCS Expellant 8.1Spacecraft Subtotal 234.9Systems Margin 20 %Spacecraft Total Mass 281.9

Table 6.10.1: Final mass budget for the low thrust transfer option.

90 7 EVALUATION AND COMPARISON OF TRANSFER METHODS

7 Evaluation and Comparison of Transfer Methods

7.1 Transfer time

The transfer time for the direct chemical Option is 5 days, while for low thrust option

it is 730.7 days, and for WSB option, it is 80 days.

This adds radiation and debris damage risks to the low thrust and the WSB option,

as scored below.

The propellant mass requirement for the direct chemical option is 100.1 kg, while it

is 96.6 kg for WSB option, and 34.3 kg for the low thrust option.

7.2 Final orbit

The ideal orbit chosen for the mission has argument of periselene (AoP) of 0◦, produced

with EMT software. Each transfer options has a different value for the final AoP,

varying from 4◦North for WSB transfer method to 40◦South for the direct chemical

transfer. These deviations from the ideal orbit could be corrected by executing ma-

noeuvres after the Moon insertion burn; but we decided to avoid it, as such burns are

expensive in terms of propellant mass. The ideal orbits will provide ability to take

sufficient picture quantity at different spatial resolutions. Hence, as a result we have

AoP of 4◦N for WSB, 40◦S for DC, and 0◦for low thrust.

Observable lunar landscape is different for the three options; achievable ground resolution

over the six-month period is shown in Figure 7.2.1.

7.3 Extended Mission Lifetime

Due to the moon’s non-spherical gravity, the final orbit affects the orbital perturbence

forces. Without correction, this would affect the rate of decay of orbit; with correction,

this affects the frequency of stationkeeping maneuvre required, hence affecting the

mission lifetime.

Though each option fulfils the mission requirement of 6 months, total duration of

time the spacecrafts are able to maintain the required orbit differ; it is desirable to

achieve longer mission life, as this would increase the ammount of data retrievable by

the mission.

The final orbit of the direct chemical option provides the least extension of the

three options, scoring 3 extra months. The WSB gives 36 months of life extension.

The longest extension is achieved by the low thrust option, due to the stable orbit

with AoP of 0◦and the high Isp of the engine.

The details of stationkeeping are described in Figures 4.1.2, 5.1.2 and 6.1.3.

7.4 Mass 91

-80 -60 -40 -20 0 20 40 60 800

20

40

60

80

100

120

South - Lunar Latitude [deg] - North

Ground Resolution [m]

Achievable Ground Resolution by the Three Options

Direct Chemical option

WSB option

Low Thrust option

Figure 7.2.1: Achievable ground resolution of each design option over the six-monthperiod

7.4 Mass

During the design optimisation stage, it was aimed to keep the final spacecraft mass

below the given 300 kg. It was found that the major contributor to the overall mass

were the main propulsion, AOCS and Power subsystems.

During the design process, it was decided that any mass left would be used to carry

stationkeeping fuel. An algorithm to do so has been embedded into the spreadsheets,

achieving launch mass of just under 300kg for direct chemical and WSB options.

For the low thrust option, however, it has been found that using this algorithm

would add tens of years’ worth of stationkeeping fuel. It has been decided that

this would not be necessary, as some other system would most likely fail before this

happens. As a result, the low thrust option weighs 282kg in total. This provides an

opportunity to accommodate further payloads, or redundancies, increasing the value

of the mission.

As a reference, it has been found that Ariane 5G charges $10000 per kg. Hence

the lighter the spacecraft is, the cheaper the launch would be.

92 7 EVALUATION AND COMPARISON OF TRANSFER METHODS

7.5 Risk Assessment

In the risk assessment table, Table 7.5.1, the X column quantifies the hazard on the

system, scaled from one to five, with one being minimal damage and five representing a

total loss of the spacecraft. The Y column is the likelihood of the hazard to occur. The

risk column shows the weighted factor applied on the systems, affected by the occurred

hazard. After considering the three options and its risk with following factors, it was

found that the direct chemical option possesses the least risk of total loss of mission,

that the low thrust option has the highest risk of the three options.

Failure mode Type of Hazard X Y Risk Redun-dancy?

Heaters Failure to maintain temperature; 5 0.1 0.5 YesCoating Penetration of solar flux into 5 0.3 1.5 Yesablation the on-board systemBattery Inability to store electric power 5 0.1 0.5 Nofailure from solar arraysSolar arrays Failure to generate solar power 4 0.3 1.2 Yeswiring failureSolar array No generation of power leading 5 0.5 2.5 Nodeployment to mission failurePCDU Failure to distribute power 3 0.4 1.2 YesMain engine Orbital insertion failure 5 0.1 0.5 NoValves Improper flow of propellant to the 3 0.6 1.8 Yes

engine causing variable thrustthen desired

Fuel Explosion and damage to 5 0.2 1 Noleakage / subsystems while posing a threatfreezing to thermal subsystemReaction Unable to move the spacecraft 3 0.1 0.3 Yeswheels along its three axisSensors Failure to perform orientation 2 0.2 0.4 Yes

around a reference pointThrusters Unable to perform Attitude 5 0.5 2.0 Yes

Control manoeuvresPayloads Unable to meet primary 4 0.1 0.4 Yes(primary) mission requirementPayloads Circuit imbalance and inability to 2 0.1 0.2 No(secondary) achieve secondary objectivesComputers Loss of control over a subsystem 5 0.3 1.5 YesDownlink Unable to send data back to 5 0.5 2.5 Yesfailure ground station causing on board

data handling crashUplink Unable to send command to the 4 0.2 0.8 Yesfailure spacecraft causing communication

malfunction and might lead tomission failure

Table 7.5.1: Assessment of system points of failure.

93

Transfer Options Major Failure modes causing Totalmission failure Risk

Direct Chemical Main Nozzle, Valves, Fuel Leakage,Sensors, 5.6Option Payloads, Computer, Communication, HeatersWeak Stability Main Nozzle, Valves, Fuel, Leakage,Sensors, 7.1Boundary Payloads, Computer, Communication, Coating

ablation, HeatersLow Thrust Option Main Nozzle, Valves, Fuel Leakage,Sensors, 10.8

Payloads, Computer, Communication, Coatingablation, Heaters, Solar arrays wiring failure,Solar array deployment

Table 7.5.2: Major failure modes for the different transfer options.

8 Observation Strategy

To achieve the primary and secondary science goals, most of the pictures and data

from other payloads are collected when the spacecraft is near its periselene. When in

aposelene, other tasks as down- and uplink communications and momentum dumping

maneuvres can be performed. Some pictures would also be taken near the aposelene,

to achieve variations in the view of the moon, as outlined in the primary science goals.

Due to the excess power available for the low thrust option (see Sections 4.3, 5.3

and 6.3), X-band downlink transmitter for the low thrust option is able to operate at

higher power, increasing the number of pictures retrievable to 17 per orbit, compared

to 10 per orbit for the other two options.

The strategies are unique for each transfer options, and are reflected in the power

profile graphs, shown in Figures 4.3.1, 5.3.1 and 6.3.1 for direct chemical, WSB and

low thrust option respectively.

Main regions of interest on the lunar surface have been identified and listed in

Table 8.0.3, and an STK simulation was performed to show that these regions would

be covered for the 94 % of the orbit.

The RAD and C1XS would remain activated on all instances of the orbit while

the camera would be switched on above the particular regions of interest. C1XS shall

perform studies at different range of spectroscopy to create a detailed map of lunar

surface for each element type.

94 8 OBSERVATION STRATEGY

Main Crater/Placeof Interest

Reason of Interest INSTRUMENT

CabeusPresence of Hydrogen gas CAMERA, C1XS84◦54’0”S 35◦30’0”W

Centaur Impact siteCentaur Impact RAD84◦54’0”S 35◦30’0”W

Apollo 12 landingsite(Oceanus Pro-cellarum)

History of lunar sedimentation CAMERA, C1XS

3◦11’51” S 23◦23’8” WApollo 14(Fra Mauroformation) Geological history ALL6◦0’0”S 17◦0’0”WCopernicus

Geographical ridges and peaks CAMERA9◦42’0”N 20◦0’0”WMare Frigoris

Iron, Titanium and water CAMERA, C1XS56◦0’0” N 1◦24’0” EAristoteles

Geographical features CAMERA50◦12’0”N 17◦24’0”EMare Imbrium

Information about lunar past ALL32◦48’0”N 15◦36’0”WHerschel

Geological features Camera,C1XS5◦42’0”S 2◦6’ 0” WMare Insularum

Uranium belt Camera, C1XS7◦60’19”N 30◦59’1 WMare Moscoviense*

Far side Thorium Belt C1XS27◦28’ N 148◦12’EMare Ingenii*

Far side Fe belt CAMERA, C1XS33◦25’S 164◦54’EApollo Landing Sites Apollo Landing Site CAMERA

Table 8.0.3: Region of interest on lunar surface

95

Figure 8.0.1: STK Simulation for Observation Strategy

96 9 FINAL CHOICE OF ORBITER

9 Final Choice of Orbiter

After having designed three different spacecrafts, one spacecraft was selected for rec-

ommendation based on the reasons described below.

Despite the advantage that the low thrust option can transmit 17 pictures per

orbit, compared to 10 for other options, the low thrust option is discounted for its

high risk imposed by the transfer. Although the transfer time itself is not a deciding

factor, the potential degradation associated with the long transfer duration (over two

years) is. Additionally, debris impacts will accumulate and could be detrimental to the

mission. Furthermore, the time spent in the Van Allen radiation belts is also higher

than for the other two transfer options, which poses further risks to the subsystems.

The second major risk specific to the low thrust option is imposed by the deployable

solar arrays. The arrays could fail to deploy leaving the spacecraft powerless, resulting

in the total loss of the mission. Both the WSB and direct chemical options have much

shorter transfer times (5 and 80 days respectively) and the solar arrays are body

mounted, which makes them power safe.

The other two options use chemical propulsion, which has much more extensive

flight heritage: it is used for the majority of space missions (including the Apollo moon

missions). As transfer times for both the WSB and the direct chemical method are

relatively low, there is little difference in the risks of degradation and debris impact.

The masses of the two spacecraft are around 300 kg, eliminating satellite mass as a

deciding factor. The final orbit for these two options does differ however, and with

it the orbit-maintenance ∆ Vs. The WSB option has a final orbit that deviates

4◦from the ideal inclination of the orbital plane (compared with direct chemical’s

40◦) and only requires 78 m/s a year for station-keeping (compared to 340 m/s for

direct chemical). This allows for a mission lifetime of 1350 days when using the WSB

transfer, which is an order of magnitude greater than the 298 day mission lifetime for

the direct chemical transfer.

The WSB route with the lunar swing-by offers the chance to demonstrate this

novel approach to lunar transfer methods. Therefore, the final recommendation is for

the orbiter that employs the weak stability boundary transfer method.

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lunar educational wide imaging satellite

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mission requirements

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mission objectives

propulsion subsystem

achemical transfer

power subsystem

direct chemical method

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