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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 Mission Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 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
5.3.4 Solar Array Sizing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 62
5.3.5 Solar Array Configuration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 62
5.3.6 Power Profile . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 62
5.4 OBDH . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 62
5.5 Communication Subsystem . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 63
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
5.7.1 Primary Assumptions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 65
5.7.2 Method of Calculations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 65
5.7.3 Equilibrium and Eclipse Temperatures . . . . . . . . . . . . . . . . . . . . . 65
5.7.4 Thermal Control Decisions . . . . . . . . . . . . . . . . . . . . . . . . . . . . 65
5.8 Structure and Configuration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 66
5.9 Discussion of the Budget Evolution due to Sub-system Trade-offs . . . . . . . . . . . 68
5.10 Final Mission Budget . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 70
6 Low Thrust Transfer Method 71
6.1 Mission Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 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
6.3.4 Solar Array Sizing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 77
6.3.5 Solar Array Configuration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 77
6.3.6 Power Profile . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 78
6.4 OBDH . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 78
6.5 Communication Subsystem . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 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
6.7.1 Primary Assumptions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 81
6.7.2 Method of Calculations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 82
6.7.3 Equilibrium and Eclipse Temperatures . . . . . . . . . . . . . . . . . . . . . 82
6.7.4 Thermal Control Decisions . . . . . . . . . . . . . . . . . . . . . . . . . . . . 82
6.8 Structure and Configuration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 83
6.9 Discussion of the Budget Evolution due to Sub-system Trade-offs . . . . . . . . . . . 86
6.10 Final Mission Budget . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 89
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
18 4 DIRECT CHEMICAL TRANSFER METHOD
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.
4.2 Chemical Propulsion Subsystem 19
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
20 4 DIRECT CHEMICAL TRANSFER METHOD
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.
4.2 Chemical Propulsion Subsystem 21
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.
22 4 DIRECT CHEMICAL TRANSFER METHOD
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]
24 4 DIRECT CHEMICAL TRANSFER METHOD
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)
4.3 Power Subsystem 25
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.
26 4 DIRECT CHEMICAL TRANSFER METHOD
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.
28 4 DIRECT CHEMICAL TRANSFER METHOD
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
4.4 On-Board Data Handling Subsystem (OBDH) 29
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.
30 4 DIRECT CHEMICAL TRANSFER METHOD
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.
4.4 On-Board Data Handling Subsystem (OBDH) 31
..
..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.
32 4 DIRECT CHEMICAL TRANSFER METHOD
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)
34 4 DIRECT CHEMICAL TRANSFER METHOD
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 Communication Subsystem 35
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.
36 4 DIRECT CHEMICAL TRANSFER METHOD
Figure 4.5.1: Zenith attenuation of signal vs frequency [42]
4.5 Communication Subsystem 37
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
38 4 DIRECT CHEMICAL TRANSFER METHOD
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.
4.5 Communication Subsystem 39
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.
40 4 DIRECT CHEMICAL TRANSFER METHOD
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
42 4 DIRECT CHEMICAL TRANSFER METHOD
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
4.6 Attitude Determination and Control Subsystem (ADCS) 43
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
44 4 DIRECT CHEMICAL TRANSFER METHOD
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 Attitude Determination and Control Subsystem (ADCS) 45
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.
46 4 DIRECT CHEMICAL TRANSFER METHOD
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.
48 4 DIRECT CHEMICAL TRANSFER METHOD
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.
4.7 Thermal Control Subsystem 49
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 ,
50 4 DIRECT CHEMICAL TRANSFER METHOD
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
4.7 Thermal Control Subsystem 51
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.
52 4 DIRECT CHEMICAL TRANSFER METHOD
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.
54 4 DIRECT CHEMICAL TRANSFER METHOD
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.
56 4 DIRECT CHEMICAL TRANSFER METHOD
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
Snapshots during development
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
Snapshots during development
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
Snapshots during development
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 Mission Analysis
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
Orbital Radius [Moon Radius]
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.
5.2 Chemical Propulsion Subsystem 59
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.
60 5 WEAK STABILITY BOUNDARY TRANSFER METHOD
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.
5.3 Power Subsystem 61
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.
62 5 WEAK STABILITY BOUNDARY TRANSFER METHOD
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.
5.3.3 Battery Sizing
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.
5.3.4 Solar Array Sizing
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.
5.3.5 Solar Array Configuration
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.
5.5 Communication Subsystem 63
5.5 Communication Subsystem
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.
Mission phase Data rate
Transfer orbit 300 bpsMission orbit (low gain) 2.03 kbpsMission orbit (high gain) 643.68 kbpsContingency data rate incase of high gain antennafailure
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.
5.6.3 Selection and Sizing of ADCS Hardware
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.
64 5 WEAK STABILITY BOUNDARY TRANSFER METHOD
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
5.7 Thermal Control Subsystem 65
5.7 Thermal Control Subsystem
5.7.1 Primary Assumptions
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.
5.7.2 Method of Calculations
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.
5.7.3 Equilibrium and Eclipse Temperatures
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
Table 5.7.1: Temperatures Reached at Each Mission Stage
5.7.4 Thermal Control Decisions
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.
66 5 WEAK STABILITY BOUNDARY TRANSFER METHOD
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.
5.8 Structure and Configuration
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:
Distance from geometric centre (cm)
Dx 2Dy -4Dz -10
Table 5.8.2: Gravity centre
The moments of inertia are then almost identical to the ones for direct chemical
and given in Table 5.8.3.
5.8 Structure and Configuration 67
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
68 5 WEAK STABILITY BOUNDARY TRANSFER METHOD
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
Snapshots during development
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
Snapshots during development
Mas
s [k
g]
mdry
mexp, trsf.
mexp, AOCS
mmargin, sys.
Figure 5.9.2: Evolution of the satellite wet mass budget during development.
70 5 WEAK STABILITY BOUNDARY TRANSFER METHOD
0
100
200
300
400
Snapshots during development
Pow
er r
equi
rem
ent [
W]
ppayload
pcomms.
pOBDH
pAOCS
ppropulsion
ppower
pthermal
pmargin, sys.
Figure 5.9.3: Evolution of the satellite power budget during development.
0
100
200
300
400
Snapshots during development
Life
time
[day
s]
tcommitioning
ttransfer
toperation
Figure 5.9.4: Evolution of the mission lifetime during development.
5.10 Final Mission Budget
Subsystem Mass Component Component AllocationMargin Subtotal
[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 Mission Analysis
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
Orbital Radius [Moon Radius]
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
74 6 LOW THRUST TRANSFER METHOD
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.
76 6 LOW THRUST TRANSFER METHOD
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 Power Subsystem 77
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.
6.3.3 Battery Sizing
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.
6.3.4 Solar Array Sizing
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.
6.3.5 Solar Array Configuration
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]
78 6 LOW THRUST TRANSFER METHOD
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.
6.5 Communication Subsystem
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.
Mission phase Data rate
Transfer orbit 300 bpsMission orbit (low gain) 2.03 kbpsMission orbit (high gain) 643.68 kbpsContingency data rate incase of high gain antennafailure
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
80 6 LOW THRUST TRANSFER METHOD
Figure 6.6.1: Disturbance Torque variation during transfer
disturbance torque during transfer is TT = 2.52× 10−6 Nm.
6.6.3 Selection and Sizing of ADCS Hardware
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
6.7 Thermal Control Subsystem 81
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
6.7 Thermal Control Subsystem
6.7.1 Primary Assumptions
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].
82 6 LOW THRUST TRANSFER METHOD
6.7.2 Method of Calculations
The method given in 4.7.2 will be used to find the equilibrium and transient tem-
peratures, and to size any heaters needed.
6.7.3 Equilibrium and Eclipse Temperatures
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
Table 6.7.2: Temperatures Reached at Each Mission Stage
6.7.4 Thermal Control Decisions
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
6.8 Structure and Configuration 83
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.
6.8 Structure and Configuration
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.
84 6 LOW THRUST TRANSFER METHOD
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.
Distance from geometric centre (cm)
Dx 0.8Dy 5Dz -5
Table 6.8.2: Gravity centre
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.
6.8 Structure and Configuration 85
Figure 6.8.1: Low thrust propulsion model, perspective top view
Figure 6.8.2: Low thrust model, perspective side view
86 6 LOW THRUST TRANSFER METHOD
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
Snapshots during development
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
Snapshots during development
Mas
s [k
g]
mdry
mexp, trsf.
mexp, AOCS
mmargin, sys.
Figure 6.9.2: Evolution of the satellite wet mass budget during development.
88 6 LOW THRUST TRANSFER METHOD
0
500
1000
1500
2000
2500
3000
Snapshots during development
Pow
er r
equi
rem
ent [
W]
ppayload
pcomms.
pOBDH
pAOCS
ppropulsion
ppower
pthermal
pmargin, sys.
Figure 6.9.3: Evolution of the satellite power budget during development.
0
100
200
300
400
500
600
700
800
900
1000
Snapshots during development
Life
time
[day
s]
tcommitioning
ttransfer
toperation
Figure 6.9.4: Evolution of the mission lifetime during development.
.
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]
6.10 Final Mission Budget
Subsystem Mass Component Component AllocationMargin Subtotal
[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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