Design study for a Rapid Imminent Impactor Characterization system

CRANFIELD UNIVERSITY
GEORGIOS GALANOS
RIICS:
Rapid Imminent Impactor Characterization System:
ELECTRICAL POWER SYSTEM
SCHOOL OF AEROSPACE, TRANSPORT AND
MANUFACTURING
Group Design Project
MSc in Astronautics and Space Engineering
Academic Year: 2018 - 2019
Supervisor: Dr Joan-Pau Sánchez Cuartielles
October 2018

CRANFIELD UNIVERSITY
SCHOOL OF AEROSPACE, TRANSPORT AND
MANUFACTURING
Group Design Project
MSc in Astronautics and Space Engineering
Academic Year 2018 - 2019
GEORGIOS GALANOS
RIICS: ELECTRICAL POWER SYSTEM
Supervisor: Dr Joan-Pau Sánchez Cuartielles
October 2018
This report is submitted in partial fulfilment of the requirements for
the degree of MSc in Astronautics and Space Engineering
© Cranfield University 2018. All rights reserved. No part of this
publication may be reproduced without the written permission of the
copyright owner.

Georgios Galanos
i
ABSTRACT
This report is part of a group project of fifteen students of Cranfield University
and is the preliminary design of the RIICS mission: Rapid Imminent Impactor
Characterisation system.
RIICS is a science-driven mission, which aims to characterise and detect near
earth objects and exoplanets as a secondary science. A low cost mission of a
50M € budget.
This report is the written proof work of Georgios Galanos, a member of the
group project and responsible for the electrical power subsystem of the
spacecraft.
This report contains an analytical design of the electrical power subsystem of
the RIICS project. The main requirement of the project is to design a low cost
and reliable system. The final design successfully achieved to fulfil the above
requirements (low cost and reliable power system).
This report assesses the final design of the electrical power subsystem. The
report includes the analysis and sizing of the primary and secondary power
source, and comparison of the different design options. It also includes the main
power control and distribution system to the loads.
Keywords: Electrical, Power, Solar P-V, Batteries, PCDU

Georgios Galanos
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ACKNOWLEDGEMENTS
First of all I would like to thank all the group members of the RIICS project and
our supervisor Dr Joan-Pau Sánchez Cuartielles who was meeting us every
single week from the beginning of the year to help and advise us. Moreover, I
would like to thank Dr Leonard Felicetti who taught me about power systems
with the most efficient way.
Finally I would like to thank my family for always being by my side.

Georgios Galanos
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TABLE OF CONTENTS
© Cranfield University 2018. All rights reserved. No part of this publication
may be reproduced without the written permission of the copyright owner......... ii
ABSTRACT ..........................................................................................................i
ACKNOWLEDGEMENTS................................................................................... iii
LIST OF FIGURES............................................................................................ vii
LIST OF TABLES ............................................................................................... ix
LIST OF ABBREVIATIONS...............................................................................xiii
1 Introduction.......................................................................................................1
1.1 RIICS Background .....................................................................................1
1.2 Electrical power background......................................................................1
2 Main power sources .........................................................................................3
2.1 Comparison of primary power sources based on mission’s
requirements....................................................................................................6
3 Power Control...................................................................................................7
3.1 Direct Energy Transfer (DET) ....................................................................8
3.1.1 Fully regulated Bus..............................................................................8
3.1.2 Sun - Regulated Bus (Unregulated Bus).............................................9
3.1.3 Comparing Regulated and Unregulated Bus.....................................10
3.2 Peak Power Tracker (PPT) - Power control.............................................11
3.3 Power Control Trade-off...........................................................................12
3.4 Bus voltage selection...............................................................................14
4 Solar arrays ....................................................................................................15
4.1 Solar PV technology ................................................................................15
4.2 Solar array structure ................................................................................16
4.2.1 Rigid Panels ......................................................................................16
4.2.2 Body–Mounted ..................................................................................17
4.2.3 Three or More Wings.........................................................................18
4.2.4 Flexible Array ....................................................................................19
4.2.5 Selection of Construction ..................................................................19
5 Energy storage - Secondary battery...............................................................20
5.1 Types of secondary batteries...................................................................21
6 Analyze and Size the Power System..............................................................23
6.1 Study Case 1 ...........................................................................................23
6.2 Study Case 2 ...........................................................................................25
6.3 Comparison of Cases ..............................................................................26
6.4 Power Consumption Analysis ..................................................................29
7 Solar Arrays Sizing.........................................................................................36
7.1 Battery’s Component Selection................................................................38

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7.1.1 Configuration.....................................................................................38
8 Power Control and Distribution Unit (PCDU)..................................................40
9 Conclusion......................................................................................................44
REFERENCES..................................................................................................45
BIBLIOGRAPHY................................................................................................47
APPENDICES ...................................................................................................49

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LIST OF FIGURES
Figure 1-1 Typical Architecture of a Power System ...........................................2
Figure 2-1 Energy sources options for various power requirements ..................6
Figure 3-1 Regulated Bus DET...........................................................................9
Figure 3-2 Sun- Regulated Bus (Unregulated Bus) DET..................................10
Figure 3-3 Peak Power Tracker Architecture....................................................11
Figure 3-4 Battery charge and discharge options in peak power tracking
architecture.................................................................................................12
Figure 3-5 Optimum voltage for various power levels ......................................14
Figure 4-1 Typical Solar Array Design Parameters ..........................................15
Figure 4-2 A fully deployed dollar array............................................................16
Figure 4-3 Panels section with cells mounted on a honeycomb substrate with
face sheets.................................................................................................17
Figure 4-4 A body mounted solar array ............................................................18
Figure 4-5 Three wings array architecture........................................................18
Figure 4-6 Final design of the Solar arrays.......................................................20
Figure 6-1 Solar arrays comparison .................................................................27
Figure 6-2 Batteries Comparison......................................................................28
Figure 6-3 Total Mass and Cost Comparison...................................................29
Figure 6-4 Average Power Consumption..........................................................34
Figure 6-5 Power Consumption during Operational Phase ..............................35
Figure B-1 Selected Orbit.................................................................................62
Figure B-3 Soyuz ST-B - SYLDA-S dual launch configuration.........................63
Figure B-4 Mission Phases Timeline................................................................64
Figure B-5 Launch and Early Operations Phase Timeline ...............................66
Figure B-6 Commissioning Phase Timeline .....................................................68
Figure B-7 Paints and surface coatings............................................................72
Figure B-8 Overview of the shapes of the secondary structures......................74
Figure B-9 Transfer phase propulsion schematic.............................................75

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Figure B-10 Reaction Control Schematic after propulsion module separation.75
Figure B-11 Fully deployed spacecraft overview in its operational configuration
at the L1 point.............................................................................................78
Figure B-12 Folded launch configuration combining the spacecraft and the
propulsion module......................................................................................78
Figure B-13 Compatibility verification with the upper position of the SYLDA-S78
Figure B-14 List of the main components of the spacecraft .............................79
Figure B-15 Outer and Inner Structure of the main spacecraft.........................79
Figure B-16 External configuration of the main spacecraft...............................80
Figure B-17 Internal bottom body configuration ...............................................80
Figure B-18 Internal top body configuration .....................................................80
Figure B-19 Overview of the propulsion module configuration.........................81
Figure B-20 Spacecraft Overview Design ........................................................82
Figure B-21 Cassegrain Design of Telescope..................................................82
Figure B-22 Probability inside scan area centered (-120,0) .............................86
Figure B-23 Position at detection comparison with limiting magnitude 21 and 17
...................................................................................................................88
Figure B-24 Difference between maximum apparent motion with perturbations
and maximum apparent motion without perturbations .............................89
Figure B-25 Mean apparent motion for each object in degrees per day........89
Figure C-1 Shunt Regulator specifications and functional schematic.............115
Figure C-2 Pyro Firing Drive Module specifications and functional schematic
.................................................................................................................116
Figure C-3 Equipment Power Distribution Module specifications and functional
schematic .................................................................................................117
Figure C-4 Battery Charge / Discharge Regulator specifications and functional
schematic .................................................................................................118
Figure C-5 Heater Power Distribution Module specifications and functional
schematic .................................................................................................119
Figure C-6 Modular Medium Power Unit specifications..................................120

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LIST OF TABLES
Table 2-1 Technology options and status ...........................................................5
Table 3-1 Advantages and Disadvantages of different architectures................13
Table 5-1 Number of cells as a factor of bus voltage........................................20
Table 5-2 Issues in Designing the Energy Storage Capability..........................21
Table 5-3 Characteristics of Selected Secondary Batteries .............................22
Table 6-1 Solar Arrays Specifications Case 1...................................................24
Table 6-2 Battery Specifications Case 1...........................................................24
Table 6-3 Total Specifications Case 1 ..............................................................25
Table 6-4 Solar Arrays Specifications Case 2...................................................26
Table 6-5 Battery Specifications Case 2...........................................................26
Table 6-6 Total Specifications Case 2 ..............................................................26
Table 6-7 Battery Discharge for cases tat peak power is required ...................31
Table 6-8 Cycle life of battery ...........................................................................32
Table 6-9 Part of the Power Break During Operational Phase .........................34
Table 6-10 Battery Specifications for operational phase...................................36
Table 7-1 Solar arrays main features................................................................38
Table 7-2 Battery’s worst Case Scenario Parameters ......................................39
Table 7-3 VES 140 Shaft Battery......................................................................40
Table B-1 Highest Criticality Events and Key Prevention Actions ....................52
Table B-2 Trade off...........................................................................................52
Table B-3 Mass Budget Breakdown .................................................................53
Table B-4 AOCS Mass Breakdown...................................................................55
Table B-5 HGA Link Budget..............................................................................56
Table B-6 LGA Link Budget ..............................................................................57
Table B-7 Power Budget...................................................................................60
Table B-8 Propellant mass and ΔV break down ...............................................60
Table B-9 Development Cost of the spacecraft ................................................61

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Table B-10 Overall Mission Cost Budget..........................................................61
Table B-11 Selected Orbit.................................................................................62
Table B-12 Orbit’s Characteristics ....................................................................62
Table B-13 Transfer trajectory parameters.......................................................63
Table B-14 Mission Phases Timeline................................................................65
Table B-15 Launch and Early Operations Phase Timeline...............................67
Table B-16 Commissioning Phase Timeline.....................................................69
Table B-17 Science Operations Priorities.........................................................69
Table B-18 Thermal Control Breakdown...........................................................71
Table B-19 Main thermal control elements .......................................................71
Table B-20 Properties of the central tubes of the primary structure .................72
Table B-21 Properties of the vertical panels of the primary structure...............73
Table B-22 Properties of the horizontal decks of the primary structure............73
Table B-23 Properties of the secondary structures...........................................74
Table B-24 Size, throughout and data produced estimates for the telescope ..76
Table B-25 Final size and throughput estimates...............................................76
Table B-26 Onboard Computer performance specifications.............................76
Table B-27 Communications design hardware selection mass and power
budgets.......................................................................................................77
Table B-28 Main Mechanisms Configurations ..................................................81
Table B-29 Optical Configuration of Telescope ................................................83
Table B-30 Integration time to achieve a SNR = 5 with 0.3m aperture.............83
Table B-31 Modified visible camera from the UVIS instrument specifications..84
Table B-32 NIR Spectrometer specifications....................................................84
Table B-33 Requirements imposed for baseline design ...................................85
Table B-34 Telescope performance analysis for secondary science operations
...................................................................................................................85
Table B-35 Requirement performance in Characterisation mode.....................86
Table B-36 Requirement performance in Scan mode.......................................86

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Table B-37 Warning time with the frozen design ..............................................87
Table C-1 Power Breakdown. Analytic Power consumption for each instrument,
charge and discharge mode and total consumption of each phase.........111
Table C-2 Analytical operation of the battery during each phase of the mission
.................................................................................................................112
Table C-3 Analytic data for the cycle life of the battery...................................113
Table C-4 Comparison between VES 100, VES 140 and VES 180.
Specifications of each battery type...........................................................114
Table C-5 PCDU Dimensions and Mass.........................................................114

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LIST OF ABBREVIATIONS
AU Astronomical Unit
BCDR Battery Charge / Discharge Regulator
BOL Beginning of Life
CM Command and Monitoring
DC Direct Current
DET Direct Energy Transfer
DoE Department of Energy
EOC End Of Charge
EOD End Of Discharge
EOL End of Life
EPD Equipment Power Distribution
ESA European Space Agency
GaAs Gallium Arsenide
GEO Geostationary Orbit
HPD Heaters Power Distribution
ISS International Space Station
LCL Latching Current Limiters
Li-ion Lithium
MJ Multi Junction
MMPU Modular Medium Power Unit
NASA National Astronautics and Space Administration
NEO Near Earth Object
Ni-Cd Nickel Cadmium
Ni-H2 Nickel Hydrogen
PCDU Power Control and Distribution Unit
PFD Pyro Firing Drive
PPT Peak Power Tracker
PV Photovoltaic
RIICS Rapid Imminent impactor Characterization
RSO Resident Space Objects

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RTG Radioisotope Thermoelectric Generators
S4R Shunt Regulation Module
Si Silicon
SJ Single Junction
SSA Space Situation Awareness
SSTL Surrey Satellite Technology
TE Thermoelectric
TJ Triple Junction

Introduction Georgios Galanos
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1 Introduction
1.1 RIICS Background
RIICS mission is based on the ESA’s SSA (Space Situation Awareness
Program). The SSA program has been created from the need of awareness to
predict and detect man-made space orbits, in-orbit events, potential impacts of
NEO’s and effects of space weather phenomena and ground based
infrastructures. As a result, the life risk and other undesired situations such as
the Chelyabinsk meteorite will be eliminated. NEOs can be defined as asteroids
or comets that pass near the Earth. SSA program aims to understand these
kinds of objects in order to decrease the risk of causing damages. Around
600,000 asteroids are known in our Solar system and 16,000 of them are
classified as NEOs. Constant and efficient monitoring has to be carried out to
ensure Earth is not being affected by potential impacts.
RIICS mission is designed to achieve two main objectives. The first objective is
the physical characterisation and scanning system for NEO’s on the order of
few meters. A better understanding of the NEO’s can be obtained by
characterising imminent impactors and comparing the information collected
prior and to the impact. Current ground telescopes have low capability detection
due to the atmosphere and the Sun effects, which can be avoided by setting the
telescope in orbit (Seurin, N., 2019).
During the 6-year mission (5 in operations), the spacecraft will operate
secondary science with the use of the existing telescope and sensors. The
observation and characterisation of exoplanets and identification of RSO
(resident space objects) form the secondary science of the mission.
1.2 Electrical power background
The power system is one of the most critical subsystems of a spacecraft. The
failure of the power system and inability to supply the required power to the
spacecraft results in the failure of the entire mission. (Fortescue and Stark,

Introduction Georgios Galanos
2
2003) It is of high importance to design a reliable and efficient power system. In
the early 80s, space agencies focused mainly on large satellites such as ISS
and manned missions to the moon. The necessity of using large satellites led
the agencies to design high power consuming systems. The American Nasa
and DoE and the European Space Agency (ESA) made extensive studies to
accomplish these requirements (SPS Concept). Many research projects were
forced to slow down or even stop due to political and technical issues,
regardless the critical threat of global warming (Landis, 2006). Despite that in
the beginning most space agencies attempted to build large spacecrafts,
nowadays space projects focus more on small satellites in order to maintain the
low cost and high efficiency (Fortescue and Stark, 2003).
The power subsystem has to follow one important rule:
Pow_direct_01 - The power subsystem must generate, distribute and control
the required power and spread it accordingly to all the subsystems during the
mission (Seurin, N., 2019)
To achieve this requirement, the design of each of the power system’s
components must be taken into sensitive and accurate consideration.
Figure 1-1 Typical Architecture of a Power System (Fortescue and Stark, 2003)

Main power sources Georgios Galanos
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2 Main power sources
The selection of the main power source of the spacecraft depends on several
parameters. The spacecraft’s configuration is one of these parameters. Weight,
size limitations, constraints set by the launch vehicle and heat dissipation
capability are some of the main variables that affect the power subsystem of the
spacecraft (Wertz et al, 2015). The main drivers for the sizing of the power
system are the lifetime of the mission, duration of each mode and respective
power consumption. As every subsystem is linked to each other, attitude control
scheme, orbital parameters, communications, payload, mechanisms, thermal
control and on board data handling are affecting the power source selection.
Payload is one of the most important subsystems that affects the power system,
as it is the one that sets the limitations for all the subsystems and consequently
the power subsystem. Finally, the environment of the mission is a very critical
parameter that has to be taken into account. Different missions require different
approaches of power sources and energy storage methods depending on the
type of orbit and distance from the sun, especially if the mission is
interplanetary.
Spacecraft’s primary power source:
• Primary batteries
• Solar PV – secondary battery
• Radioisotope – Thermoelectric Generators (RTGs)
• Fuel cells
• Solar Concentrator – Dynamic
• Chemical Dynamic
1. Primary batteries are producing direct current by electrochemistry. One
of their main advantage is that they are the most economical primary
source for small spacecrafts with short lifetime (Miller, Keesee, 2003).
2. Solar photovoltaic power source is the most common source for
spacecraft power systems as it can provide power in tens of watts to

4
several of kilowatts up to 20 years lifetime. This method is converging the
sun radiation power into electrical power. Hence, it is considered to be
one of the most reliable and economical primary sources. In cases where
the spacecraft is under eclipses, the spacecraft must provide energy
from a different source due to the lack of solar power. During eclipses
secondary batteries must be applied for providing the necessary power
requirements. Secondary batteries are not used only for eclipses but also
for emergency cases and peak power requirements for when a direct
transfer power is in use (Jensenh, 2003).
3. As previously noted, the type of the mission is one of the main
parameters taken into account. RTGs are used generally for
interplanetary missions, particularly in deep space, where the power
consumption is very large. One of the main advantages of RTGs is its
capability of generating power in the absence of the sun and can last up
to several decades (The Viking landers were operating for 4-6 years
supplied by RTGs). Moreover, it is insensitive to the cold of the deep
space and can be exposed to the high radiation space fields. More power
can be supplied proportionally to the spacecraft’s mass. No moving parts
and absent of fluids, safe and flight-proven, and free of maintenance are
some extra advantages of the RTGs that make that power source very
reliable. On the other hand, the fact that RTGs cannot be turned on and
off and the power is decreasing exponentially with time makes it
undesired for many types of missions. From the thermal control point of
view, RTGs must be under cooling mechanisms and coverage during the
course of the mission. The main disadvantages of the RTGs power
source are the limited conversion efficiency (5%) and high cost. (Miller,
Keesee, 2003)
4. Fuel cells are extremely flexible. They can provide power during sunlight
and eclipse. Fuel cells have a high energy density, which causes them to
be a very compact comparing solution, especially regarding solar PV.
The main disadvantage of using fuel cells is the spaceship’s required fuel

5
carriage capacity. It is a good primary source for manned mission (Wertz
et al, 2015).
5. Dynamic and chemical power sources are to be applied in future
missions (Wertz et al, 2015).
The table below provides an overview of the available technology options and
their status.
Table 2-1 Technology options and status (Patel, 2005)

6
2.1 Comparison of primary power sources based on mission’s
requirements
The selection of the primary power source must be based on the mission’s
requirements. For the RIICS mission the main requirements that affect the
power source selection are the following:
• Cost
• Orbit
• Lifetime
• Power consumption
The RIICS spacecraft is a small-sized spacecraft, which requires some
hundreds of Watts to operate. Its lifetime is estimated to be 6 years with the
option of extension, if possible. The figure below is the main guide for the
decision of the main power source.
Figure 2-1 Energy sources options for various power requirements (Angrist, 1982)

Power Control Georgios Galanos
7
According to the requirements of the mission, fuel cell, Radioisotope – TE and
solar PV with secondary batteries appear to be the most suitable approaches to
our mission.
Fuel cells and RTG have specific cost in the order of tens of thousands of $/W.
In the case of PV technology the cost ranges between 300 and 900 $/W (Wertz
and Larson, 1999). Solar dynamic systems have a cost range between 1000
and 2000 $/W, and are designed to provide much more power (Patel, 2005).
The combination of PV cells and secondary batteries is the most common and
safe method of power supply for missions orbiting the Earth or orbiting the L1
point. Overall, it may be said that in order of cost, reliability and simplicity solar
PV cells and secondary batteries are chosen as primary and secondary power
source (Johnson, 2012).
3 Power Control
“Bus voltage level, power generation and energy storage must be jointly
selected to optimize the total power” (Patel, 2005). Patel states that the design
of an Electrical Power System is very complex because of the linkage between
the system components. The size of the solar panels, batteries, general
architecture of the system, bus voltage and system components discussed
below are interacting with each other in such manner that many iterations must
be made in order to achieve the desired result. It will be noticed that throughout
this report, system components are being mentioned that are only being
discussed in later chapters.
The voltage source of a spacecraft must provide the required amount of voltage
for each load. Spacecraft’s loads often require different amounts of voltage than
the voltage amount the bus operates with (Wertz et al, 2015). DC-DC voltage
converters are generally used to control and adjust the voltage amounts.
Considering that all power requirements of the subsystems and instruments are
known, a DC-DC converter is able to spread the voltage to the subsystems as

8
required. A DC-DC converter can also maintain the voltage, which is specified
by the load, within a range.
The bus must be able to control the electrical power for all subsystems and
instruments so to prevent overcharging the battery or overheating other
subsystems or instruments, including the electrical system. To control the power
bus two techniques are usually being applied: direct energy transfer (DET) and
peak power tracker (PPT) (Wertz et al, 2015). The main difference between the
two technics is the system’s reaction to the power generated by the solar
panels.
3.1 Direct Energy Transfer (DET)
In the case of DET, the power input deriving from the solar panels is being
transferred directly to the bus. As a result, all the subsystems must run by using
the same power used by the bus. In the case that a subsystem requires more
power than the available, the battery may provide the extra required power.
A shunt regulator is usually applied to the DET system in order to increase its
efficiency and reliability (Wertz et al, 2015). The shunt regulator operates in
parallel to the solar arrays and keeps the current of the arrays away from the
subsystems and battery when it is not needed. A DET - shunt regulator system
provides high efficiency at the EOL, low mass system and cost efficiency.
However, one disadvantage is that the DET system cannot operate peak power
requirements without the power support of the battery.
The DET system can be divided into two categories: fully regulated bus and
sunlight regulated bus (often referred as unregulated bus) (Patel, 2005).
3.1.1 Fully regulated Bus
Fully regulated bus, which is also known as regulated bus, controls the bus
voltage within a range of ±2 to 5% of the nominal voltage during the entire
mission (Patel, 2005). The following figure describes the architecture of a

9
regulated bus. This type of system allows the batteries to be used in parallel
with the solar arrays, which improves the system’s reliability.
Figure 3-1 Regulated Bus DET (Patel, 2005)
3.1.2 Sun - Regulated Bus (Unregulated Bus)
To minimise the complexity of the spacecraft and the power system it is more
convenient to distribute power from both sources (solar panels and secondary
batteries) but not in parallel. The case of a direct energy transfer to the bus is
known us sun–regulated bus or unregulated bus. In this method the bus voltage
is regulated during daytime via the shunt control and is unregulated during
night-time. The basic difference between a regulated and an unregulated bus is
found on the Power Regulated Unit. In the unregulated bus, the battery charger
regulator controls the battery during daytime, however, a discharge converter is
not included in the architecture, which leads to battery discharges during night-
time via a diode ‘d’ called ‘battery discharge diode’. The battery disconnects
from the bus during day-time while it is been regulated by the shunt controller.
The system allows only the battery to be discharged at night-time and blocks
any uncontrolled charge current received from the battery (Patel, 2005). The
basic architecture of such system is the following:

10
Figure 3-2 Sun- Regulated Bus (Unregulated Bus) DET (Patel, 2005)
3.1.3 Comparing Regulated and Unregulated Bus
The power system must take into consideration both day–time and night–time
phases of the mission.
A full-regulated bus is usually applied in GEO orbits where the power
requirements are above 3kW, whereas the unregulated bus is mostly used in
satellites with requirements below 3kW. (Patel, 2005)
The sun–regulated bus is often less complex and reliable than the fully
regulated bus. The main disadvantage is that the battery cannot be used during
daytime as it is disconnected from the bus. The fact that a battery charger
converter is missing contributes in reducing the system’s cost and power.
On a fully regulated bus operation, the battery can be used during daytime. The
attendance and use of a battery when peak power is required, leads to the
reduction of the area, mass and cost of the solar arrays and could save up to
20% of the mass and area of the solar panels. By using a fully regulated bus the
mission automatically becomes more reliable as the bus is continuously

11
regulated and can provide the required voltage to the loads at any time.
Nowadays, most of the space missions are using regulated buses.
3.2 Peak Power Tracker (PPT) - Power control
It is known that solar arrays generate more power at the BOL and during cold
phases of the mission at higher voltage rates. The maximum power of the
system occurs at a point where, the power is transferred by the solar arrays at
the maximum power and operates at maximum efficiency (Jiang et al, 2002).
The maximum power point of the solar arrays can be varied with time, solar
radiation, temperature and lifetime of the solar panels (Huynh and Cho, 1999).
A suitable switching regulator must be located between the solar arrays and the
bus to control the maximum power voltage that the solar arrays produce and the
voltage that the loads need to be supplied with. The figure below displays the
main architecture of a Peak Power Tracker system.
Figure 3-3 Peak Power Tracker Architecture (Patel, 2005)
The series-switching regulator stays constant at the maximum power producing
voltage using the peak power tracker. The output voltage can be adjusted to the
required level by varying the duty ratio controller. The peak power tracker can
be activated while the battery is being charged. If not, the power that left in the
solar arrays can increase their temperature.

12
PPT is mostly used for satellites that are not able to point continuously the sun
or the solar radiation and in cases when the temperature varies at a high range.
The PPT system can be designed in three ways: series, parallel and series–
parallel as shown in the figure below.
Figure 3-4 Battery charge and discharge options in peak power tracking architecture.
(Patel, 2005)
3.3 Power Control Trade-off
The table below indicates the pros and cons of each option. Cost, mass and
efficiency are the most important variables in every mission. The table shows
the best application for each case, which is not always valid since parameters
other than the cost, mass and efficiency may be equally important.

13
Table 3-1 Advantages and Disadvantages of different architectures (Patel, 2005)
According to past missions, any architecture is suitable for specific missions,
however it is proved that this is not always true. For small satellites, which
require less than 500W of power and have low orbits, PPT is the most suitable
solution. Between 1000W and 3000W, sun-regulated bus is the most
advantageous option. For power requirements above 3000W a regulated bus is
the best approach.
At this stage a PPT system seems to be the most matching system for a
spacecraft that demands power below 500W. The fact that the spacecraft will
orbit the L1 point may add some extra things to notice before the decision.
In the L1 point the spacecraft has been designed to face the Sun during the
whole mission at the most desirable angle between the Sun and the solar
arrays to avoid high losses. It is remarkable to notice that in the L1 point the

14
spacecraft will never come across eclipses, which leads to the use of the solar
arrays power during the entire mission.
Despite that in the RIICS project the power requirements remain constant in the
operation phase and peak power is absent, in some cases the battery will be
used to split the power consumption (will be discussed later on). A DET
regulated system seems to be closer to our needs, as it is more reliable and
cost-effective.
3.4 Bus voltage selection
The selection of the bus voltage of the spacecraft must be linked with the power
requirement, nominal voltage of each of the loads and buses available in the
market, in order to reduce the cost. The figure below indicates that for the RIICS
project, in which the power requirement is less than 500W, a 28V bus is the
ideal solution. All the subsystems must comply with the voltage requirement of
the electrical power subsystem.
Figure 3-5 Optimum voltage for various power levels (Patel, 2005)

Solar arrays Georgios Galanos
15
4 Solar arrays
4.1 Solar PV technology
Nowadays there are plenty of choices in Solar PV technology. This report
examines the most common technologies in space applications: Silicon, Gallium
Arsenide (GaAs) single junction, Gallium Arsenide (GaAs) multijunction and
Gallium Arsenide (GaAs) triple junction. (De Luca, 2011)
Figure 4-1 Typical Solar Array Design Parameters (Wertz et al, 2015)
The chart above clearly shows that Silicon technology is the cheapest of all the
available technologies. As the power requirements increase so do the area and
mass of the spacecraft. According to the configuration of the spacecraft and the
launcher, the area of the solar arrays is not as significant as the reduction of the
cost of each component (Jurian, T., 2019). RIICS spacecraft will consume an
average of 222,1W power. The latter, means that the Silicon option costs
around 126,000$, and for the GaAs (SJ), GaAs (MJ), and GaAs (3J) the costs
are 283,000$, 232,000$ and 206,000$ respectively. Silicon technology is the
most inexpensive solution for the spacecraft. As mentioned previously, this
project aims to minimise the total cost of the mission and for that reason Silicon
0
5
10
15
20
25
30
35
Area
(m2)
Weight
(kg)
Cost ($)
x 10^4
Power
BOL (W)
x 10
Power
EOL (W)
x 10
Solar array type Trade - off
Silicon
GaAs (SJ)
GaAS (MJ)
GaAs (TJ)

16
technology was chosen. The performance of the power system can be adapted
and provide the required power at the Beginning of Life and End of Life at the
same level of efficiency.
4.2 Solar array structure
The configuration of the solar arrays can be divided into planar and
concentrator (Wertz et al, 2015) and each type can be divided in body or panel
mounted. Up to date, results from past missions, show that most of the satellites
use planar arrays.
4.2.1 Rigid Panels
A traditional way to build the solar array is to mount the cells onto a rigid
substrate often made from aluminium and carbon face sheets. Solar cell
insulation sheets like Kapton®
, Kevlar®
and fiberglass are able to successfully
reduce the mass of the array (Wertz et al, 2015). Cover glass such as fused
silica Microsheet ®
is used to protect the array from the space environment. To
successfully produce more output power by the cell, an antireflective coating is
installed so to minimise the light reflection and allow the sunlight energy to be
absorbed by the solar cells. A coating that controls the temperature of the
surface must cover the back site of the array (Patel, 2005). The following figure
displays one wing of a rigid panel solar array.
Figure 4-2 A fully deployed dollar array (Patel, 2005)

17
The solar arrays are stowed with the satellite structure in the launch vehicle until
the phase of the separation when the solar arrays are deployed. This particular
method is discussed in the mechanism’s report (Liébana Moradillo, O., 2019).
Figure 4-3 Panels section with cells mounted on a honeycomb substrate with face
sheets (Patel, 2005).
4.2.2 Body–Mounted
Body-mounted planar cells are typically used on spinning spacecrafts. During its
rotation, the spacecraft’s surface can capture the energy of the Sun. The solar
cells can increase their thermal energy, however, the constant rotation of the
spacecraft does not allow the temperature increase of the spacecraft’s
elements. Body-mounted cells operate in less efficiency compared to deployed
solar cells, as they have to run in higher temperatures. The main disadvantage
of body-mounted cells is that the area of the solar arrays will be increased
because the cells are not illuminating all the time. Consequently, the overall
cost increases, a fact that this project attempts to eliminate. RIICS spacecraft is
a 3-axis stabiliser spacecraft that needs to be stable to operate its science
functions, thus it is more convenient to use panel mounted solar arrays (Wertz
et al, 2015).

18
Figure 4-4 A body mounted solar array (Patel, 2005)
4.2.3 Three or More Wings
This solar array construction is generally used for small science mission
satellites with Peak Power Tracker. It practically offers the same benefits as a
body-mounted array, for the reason that the arrays do not always face the sun.
The benefit of using three or more wings is that the solar arrays do not interfere
with any of the instruments located on the spacecraft body. The three or more
wings construction needs different Peak Power Tracker for each of the arrays,
or else the arrays will obstruct each other (Patel, 2005).
Figure 4-5 Three wings array architecture (Patel, 2005)

19
4.2.4 Flexible Array
Hubble Space Telescope and ISS are the most well-known satellites to use a
flexible array construction. As in the rigid panel method, the arrays are stowed
on the spacecraft when the spacecraft is in the launch vehicle and rolled out or
deployed like an accordion panel after the spacecraft’s separation with the
vehicle. As all the constructions, protection from thermal risks and control of
temperature differences when exiting the eclipses must be assured. In cases
where a flexible array is not totally flat, a reduction of the power output of the
array could occur. A temperature difference reshapes the array. Once the
temperature is being equalised, which should happen approximately 30 minutes
after the temperature changes, the array returns to its normal flat shape. This
inefficiency of the solar arrays is something that a rigid panel will never face as
it transfers the front heat to the back in a very quick way (Patel, 2005).
4.2.5 Selection of Construction
There is one more construction called concentrator. Concentrator construction
uses mirrors and lenses to collect more sun light in order to generate more
power. This construction adds more complexity to the system, therefore it has
not been taken into consideration (Wertz et al, 2015).
The RIICS spacecraft is designed be a 3-axis stabilised satellite. This
information tends to discharge the possibility of having a body-mounted array,
since it is commonly used for a spinning spacecraft (Wertz et al, 2015).
Moreover, the instruments and payload of the body mounted array spacecraft
will face very high temperatures and add complexity to the thermal design.
Finally, the spacecraft in this project uses its outer surface to accommodate
instruments, as a result there is no available surface area to provide the
required power.
Rigid panels is the selected solar construction as it is the most common, reliable
and less complex system. It is also used on 3 axis-stabilised spacecrafts. It is
important to notice that in the configuration of the spacecraft, the solar arrays

Energy storage - Secondary battery Georgios Galanos
20
are able to track continuously the sun but with a worst-case scenario incidence
angle to be accounted (will be discussed later on the report). Two rigid arrays
consisted of 3 panels each, form the solar panel system. The figure below
indicates the existing design of the solar arrays.
Figure 4-6 Final design of the Solar arrays (Jurian, T.,2019)
5 Energy storage - Secondary battery
Secondary batteries are commonly used during times of eclipses and peak
power requirements for energy storage. As mentioned previously, this project
does not call for an eclipse, given that the spacecraft will orbit the L1 point. A
secondary battery will be used to provide enough power in case of emergency
and Peak Power in the course of the mission. The power system size and its
various design options will be thoroughly discussed later in the report.
Individual cells connected in series or parallel create a pack of batteries. The
number of the individual cells required depends on the bus voltage
requirements and can be defined by the following table.
Table 5-1 Number of cells as a factor of bus voltage (Samina Asif,2008)

21
The decision of the number of cells per battery is made after the analysis of the
available battery types.
There are some basic parameters that should be considered prior to the design
of the battery and the type selection. The battery must always provide constant
voltage when required. Table 5-2 displays the issues in designing a secondary
battery.
Table 5-2 Issues in Designing the Energy Storage Capability (Wertz et al, 2015)
The conversion from chemical into electrical power and reverse is done by the
battery. A secondary battery can perform these operations up to thousand times
during its lifetime. The selection of a secondary battery is basically done by
taking into account the capacity of the battery, cost, weight, cycle life
(depending on the depth of discharge) and the way the spacecraft is going to
use it. In this project the secondary battery must be of low-cost and have long
cycle life capability in order to provide enough power throughout the mission.
5.1 Types of secondary batteries
Nowadays spacecrafts use mostly three types of secondary batteries: Nickel-
Cadmium (Ni-Cd), Nickel-Hydrogen (Ni-H2) and Lithium Ion (Li-Ion) (Broussely
Pistoia and Knovel, 2007).
Up to the 80’s, Nickel-Cadmium battery was the traditional battery for 28V
spacecrafts and buses. It usually consisted of 22-23 series-connected cells.
Their nominal capacity ranges between 5 and 100Ah. Today, Nickel-Cadmium
Physical Size, weight, configuration, operating position, static and
dynamic environments
Electrical Voltage, current loading, duty cycles, number of duty cycles,
activation time and storage time, limits on depth of
discharge, and short-circuit (fault) recovery
Programmatic Cost, shelf and cycle life, mission, reliability, maintainability,
produceability and safety

22
batteries’ application is considered to be extensive and very low risk – related to
the storage system missions (Wertz et al, 2015).
After the Nickel-Cadmium battery was introduced to the industry, Nickel-
Hydrogen (Ni-H2) batteries became the most common energy storage. Ni-H2
design configuration can be divided into three categories: single pressure
vessel, common pressure vessel and individual pressure vessel. The main
difference between each of the categories of a Ni-H2 battery is the diameter of
the cells and their working terminal voltage (Wertz et al, 2015).
Li-Ion batteries have significant advantages over Ni-Cd and Ni-H2 batteries. This
kind of battery technology offers a reduction in size, higher energy density,
higher efficiency, less complexity, less costly thermal control system and lower
self-discharge rate. The main disadvantage of the Lithium Ion batteries is that
they are are approximately double the cost of Ni-Cd and Ni-H2 (Patel, 2005).
The main characteristics of each type of batteries are shown on the table below.
Battery Type Ni-Cd Ni-H2 Li-ion
Energy Density of Battery
(Whr/kg)
30 60 125
Cycle life (80% DoD) 750 500 1500
Self discharge (per month)
*100%
0.5 0.3 0.05
Cell Voltage (V) 1.2 1.2 - 3 3.7
Charge temperature
(Celsius)
0 - 40 -20 - 30 10 - 25
Discharge temperature
(Celsius)
-20 - 65 -20 -65 -20 - 60
Maintenance requirement Full discharge
every 90 days
when in full use
Full discharge
every 90 days
when in full use
Free
Table 5-3 Characteristics of Selected Secondary Batteries (Battery University, 2019)
As stated earlier, the battery used in this project must be not only cost-effective
but also capable of providing enough power to the loads. The Nickel-Cadmium
battery seems to be the ideal solution as it is reliable and economical. The only
disadvantages of the Nickel–Cadmium battery and its cycle of life and mass.

Analyze and Size the Power System Georgios Galanos
23
The RIICS spacecraft must be able to provide a charge/discharge mode for
around 1800 cycles through the mission. It is obvious that if Nickel batteries
were used, the system would require more than 4 batteries (one additional for
redundancy), which would lead to higher battery system mass and possibly
higher launching cost. Moreover, the battery must be able to be regulated from
the PCDU. As it will be explained in the following chapters, the only available
charge/discharge regulator provided by TERMA regulates only Lithium battery
systems. In addition to all the above, Clyde Space (2010) states: “Li-Ion is fast
becoming the main energy storage technology used in space applications (Li-
ion is used on ESA deep space missions Rosetta, RoLand, Mars Express, and
many more missions including all of SSTL's recent small satellite missions)“. To
conclude, a Lithium Battery is the best approach in this particular project,
considering the cycle life of the mission, high cost due to a higher launch mass,
and charge/discharge regulator factor.
6 Analyze and Size the Power System
The power subsystem is generally quite complex. It is often interacting with the
rest of the subsystems. This system is so vulnerable that each iteration could
create undesired system alternations. This chapter examines the two most
feasible and reliable Power System designs. The effort to minimise the cost,
mass and area is not always leading to a single solution. Consequently, a trade-
off must be applied.
The only constraint in deciding the final design of the Power system is the
minimisation of costs, which is the main objective of this study.
6.1 Study Case 1
After collecting the inputs of all the subsystems, the study proceeds with the first
iteration of the design. The analysis of the power requirements of each of the
subsystems proves that the design is simple and of low cost. The power
requirements for each of the subsystems remain stable only in the operational

24
modes. This conclusion means that there is no need for Peak Power or battery
use in these modes. For that reason the Power system could be designed in
such way that the solar arrays could supply the required energy to the loads at
any time of the mission without the use of batteries. This design will occur by
increasing the solar arrays area to provide enough power during the BOL and
EOL. The battery should be used only in emergencies, resulting in adding only
one pack of batteries to the system. All types of emergencies must be studied
and included in the system design. The RIICS mission is a low cost mission of
as high reliability as possible. For that reason, in the event of both solar arrays
failing to provide the required power, the satellite will still operate normally. The
satellite will automatically switch to survival mode, in which all the subsystems
will operate the minimum number of actions possible. The survival mode has
been designed to provide enough power to all the subsystems to survive for 24
hours or until the issue has been resolved. In the event of different emergency
scenarios, the duration of the mode can be extended, given that there is no limit
of power. The solar arrays will provide the necessary power. The specifications
of this design are shown on the following tables.
Solar arrays Case 1
Area (m2
) 4.6
Cost ($) 187,200
Mass (kg) 10.5
Power available (W) 424
Table 6-1 Solar Arrays Specifications Case 1
Table 6-2 Battery Specifications Case 1
Battery Case 1
Mass (k) 7.9
Cost ($) 2,374
Capacity (Whr) 989

25
Total Case 1
Mass (k) 18.4
Cost ($) 189,574
Table 6-3 Total Specifications Case 1
Although case study 1 is simple to design and it is of low mass, its relatively
high cost does not make it a good enough case for such missions.
6.2 Study Case 2
After working on the first design of the Power system of the spacecraft, the next
step is to investigate the possibility of further system improvements. Improving
the cost factor is essential for a small consuming spacecraft. An area of 4,6 m2
solar arrays appears to be large and costly compared to similar low power
consuming spacecrafts. The maximum power that the subsystem requires is
about 330W and with a direct energy transfer system the available power has to
be 412W. These amounts derive from the need in generating power from the
solar arrays during the mission. During the phases from the start-up mode until
the spacecraft reach the L1 point and starts the operation mode, there is need
of peak power. Consequently, the power system will oversized in order to
provide the required power. All these factors increase the final area required. In
order to minimize the space used by the solar arrays on the spacecraft, is best
to make use of the battery for peak power demands. As previously discussed,
peak power is required only in phases prior to the mission’s operational phase.
It also adds complexity to the design. The use of battery for peak power will add
one more component to the power system, as two batteries will require (one
extra for redundancy). According to this design, the spacecraft is already using
a battery for phases before the operational phase and for that reason a study
was carried out as for the use of the battery for splitting the power requirements
for the operational phase and decreasing even more the space used by the
solar arrays. In this design, the mass will be higher given the need for 3 battery

26
packs because of the length of the mission and the need of redundancy. The
main specifications of this case study are included in the following tables.
Solar arrays Case 2
Area (m2
) 3.06
Cost ($) 125,931
Mass (kg) 7.1
Power available (W) 278
Table 6-4 Solar Arrays Specifications Case 2
Battery Case 2
Mass (k) 23.73
Cost ($) 7,120
Capacity (Whr) 989
Table 6-5 Battery Specifications Case 2
Total Case 2
Mass (k) 30.8
Cost ($) 133,051
Table 6-6 Total Specifications Case 2
6.3 Comparison of Cases
Understanding the requirements of the mission is of utmost importance in the
process of selecting one of the above case studies. The first case study
includes the use of solar arrays and a secondary battery in the event of an
emergency. Installing such system would only reduce the weight of the
spacecraft, as the cost and the area of the solar arrays will be very high. On the
other hand, the second case of using both solar arrays and batteries for
providing power would reduce the cost and the total area of the system but not
the system’s mass. The graphs below indicate the most important differences
between the two options.
Note: The charts below include a scaled method for comparing the values.

27
Figure 6-1 Solar arrays comparison
In the comparison of the two study cases in regards to the solar arrays design, it
can be seen that based on all specifications the second case study is more
efficient. Case study 1 calls for a larger solar array to cover the power demand
of the mission. It is essential to underline that in the first case study the
unexploited power will be dissipated and put more pressure on the thermal
subsystem. In the second case study however, the dissipated power from the
solar arrays is used directly to charge the batteries. In addition, a Maximum
Power Tracker (MPP) system could possibly be added in the first case study to
ensure constant maximum power availability. This observation leads to a small
reduction of the power available, as the efficiency of a MPP system is slightly
higher than a DET system. As a result, the area of the solar arrays could be
significantly reduced, which makes a third study focusing on a MPP system
unnecessary. For all the above, case study 2 is a better approach regarding the
design of the solar arrays, especially considering the cost minimization factor. A
discussion on the battery and total compression is following, which will give a
better understanding of the two studies.
0
2
4
6
8
10
12
14
16
18
20
Area (m^2) Cost ($)
x 10^4
Mass (kg) Power
available (W)
x 10^2
Case 1 4.6 18.72 10.5 4.24
Case 2 3.06 12.59 7.1 2.78
Axis Title
Solar arrays

28
Figure 6-2 Batteries Comparison
As for the battery design, case study 1 is the ideal solution. Figure 6-2
Batteries Comparison proves that the mass and cost of the first case study
are significantly lower than the second case study. It is important though to
note, that in both cases a Lithium battery is been used, as already discussed in
Chapter 5.1. Using batteries in the cases of peak power and power split
requirements for a 6-year lifetime mission, would increase the required quantity
of the batteries to complete the mission in a reliable and safe manner.
0
1
2
3
4
5
6
7
8
9
10
Mass (kg) x 10 Cost ($) x 10^3 Capacity (Whr)
x 10^2
Case 1 0.79 2.3735 9.89
Case 2 2.373 7.12 9.89
Axis Title
Battery

29
Figure 6-3 Total Mass and Cost Comparison
The final step for the selection of the ideal design approach is to compare the
results of the most important parameters of the designs, mass and cost of the
power system. The above diagram shows the inversed results in the two cases,
a fact that adds complexity in making the final decision. This study aims on the
minimisation of the cost, which simplifies the decision process. After examining
the launch and configuration system (Jurian, T., 2019), (Picavez C., 2019), the
mass of the power system shows only a 12kg difference, which is not a
significant factor in the decision process. On the other hand, the area utilized by
the solar arrays has to be as small as possible. For these reasons, case 2 has
been selected as the final design of the power system. The sizing and analysis
of the power system design is discussed in Chapter (7).
6.4 Power Consumption Analysis
The most important requirement for the power system is to be able to generate,
control, distribute and store the energy and provide the appropriate power to all
components during the mission. To examine if this requirement is applied to our
mission, the study conducts a power breakdown. Analysing the power
0
5
10
15
20
25
30
35
Mass (kg) Cost ($) x10^4
Total
Case 1
Case 2

30
breakdown derives a better overview of the design selection and process that
was followed to complete it.
The first step is to discuss the process followed to size the battery and solar
arrays. By having the results and the power breakdown from the first simple
design, which was exclusively based on the use of solar arrays as a power
source, the study attempts to reduce the solar array space. Table 6-1, Table 6-2
and Table 6-3 provide the basic specifications of the first power system’s design.
The battery has been sized at first for emergencies. After the selection of the
second case, in which the battery supports the system in peak power and
power split requirements during the operational phase, the power availability of
the existing battery in the event of peak power has to be examined. As it can be
seen from the table below, the total energy battery discharge in any phase is
much lower than the total energy of the battery (988Whr), therefore there is no
need to oversize the battery.
Phase Subsystem Energy
Discharged
(Whr)
Total Battery
Discharge (W)
Orbit Transfer Payload 221.1
Structure and
Mechanics
Thermal
Comms 61.2
OBDH
ADCS
Propulsion 160.6
Sun Safe During
Commissioning
Payload 169.35
Structure and
Mechanics
1.4
Thermal
Comms 150.25
OBDH
ADCS 12.70
Propulsion 5

31
Start up Payload 49.44
Structure and
Mechanics
1.34
Thermal
Comms
OBDH
ADCS
Propulsion 48.1
Stand By Payload 8.4
Structure and
Mechanics
Thermal
Comms
OBDH
ADCS
Propulsion 8.4
Table 6-7 Battery Discharge for cases tat peak power is required
A more analytical document can be found in Appendix (9C.2). The energy each
phase needs is much lower than the battery’s, which allows the operation of
more than one cycles prior to the battery charge. A study was made to calculate
the number of battery cycles that each phase can operate with in order to
extend the battery’s lifespan. The following table states the numbers concluded
by this study.
Phase Orbit transfer Sun Safe (During
commissioning)
Total Capacity of Battery
(Whr)
988.95 988.95
Capacity that needs
(Whr)
277.26 211.65
Times that battery can
be used
3.6 4.7
Duration of the Phase 173 (days) 30 (days)
Cycles of Battery 49 7

32
Phase Start-Up Stand-By
Total Capacity of
Battery (Whr)
988.95 988.95
Capacity that needs
(Whr)
61.7675 0.0105
Times that battery can
be used
16. 94185.7
Duration of the Phase 17.217 (hours) 1 (hour)
Cycles of Battery 1 1
Table 6-8 Cycle life of battery
During the research for the selection of the battery, which is discussed later on,
there were no available products in the market with an acceptable battery cycle
life and for this reason a factor of 80% of Depth of Discharge was taken into
account to provide around 1500 cycles (Battery University, 2019). It is of high
importance to use the battery efficiently and reduce the battery life cycles and
battery packs, whilst targeting the extension of the mission.
The next table displays the power consumption, charge/discharge mode,
energy consumed by the battery and power required for charging the battery
during the operational phase. A more detailed power breakdown document can
be found in Appendix 9C.1. The detailed power breakdown document contains
information about the power consumption and duration of each instrument of
the subsystems.
Phase Sub-
system
Aver.
Power
(W)
Peak
Power
(W)
Charge Dis-
charge
Energy
Disch.
(Whr)
Power
to
charge
(W)
Total
Power
(W)
Scan. Payload 56 OFF ON 222.1
Structure
/
Mechanic
s
18
Thermal 44
Comms 13.5
OBDH 33
ADCS 57.6 69 136.67
Propulsio
n
0

33
Chara.
/
NEOs
Payload 56 OFF ON 222.1
Structure
/
Mechanic
s
18
Thermal 44
Comms 13.5
OBDH 33
ADCS 57.6 69 87.74
Propulsio
n
0
Target
Acqui.
Payload 0 ON OFF 49.6 222.1
Structure
/
Mechanic
s
18
Thermal 44
Comms 13.5
OBDH 28
ADCS 69
Propulsio
n
0
Orbit
Maint.
Payload 0 ON OFF 85.54 219.1
Structure
/
Mechanic
s
18
Thermal 0
Comms 13.5
OBDH 28
ADCS 19.5
Propulsio
n
54.56
Comm Payload 0 ON OFF 137.1 221.4
Structure
/
Mechanic
s
18
Thermal 0
Comms 30.01
OBDH 33
ADCS 3.25
Propulsio
n
0

34
Table 6-9 Part of the Power Break During Operational Phase
The above table proves that there is no peak power in any of the phases except
for scanning and characterization/NEOs. This conflicts previous discussions
according to which there is no peak power in the operational phase. This
happens because on scanning and characterization/NEOs the battery shares
the power requirements of the ADCS to further decrease the solar arrays area.
Usually, the battery can be charged when not used by the system in the various
phases. In this project the battery is being used for the period of the operational
phases and modes. For that reason the battery must be charge in any phase
with available time and power for charging. Orbit maintenance, target
acquisition and communications have a total of 160 minutes available for
charging the battery.
The use of battery to split the power takes place in the operational phase. While
the battery is being used for peak power for all the other phases, operational
phase requires the maximum average power. Next chart displays the average
power for each phase of the mission.
Figure 6-4 Average Power Consumption
0.00
50.00
100.00
150.00
200.00
250.00
Average Power Consump6on
Power (W)

35
Based on the above chart, the battery is being used to split the power usage in
order to reduce the solar arrays area and cost. The maximum power that can be
supplied from the battery to the system during the operational phase has to be
examined. During operational phase the modes that consume the most average
power are the Scanning and Characterization/NEOs modes as it can be seen
on the chart below.
Figure 6-5 Power Consumption during Operational Phase
To examine how much power the battery can supply to these two modes in
every cycle of the mission, the study looks on the maximum time and power
available to the rest of the modes. The following table indicates the power
discharge for each mode, the available energy for the battery charge and the
power needed to charge the battery enough to secure the required amount of
energy.
Operational Mode
Battery
discharged
(Whr)
Energy
Available
to charge
(Whr)
Power
Available
(W)
Power
needed to
charge (W)
Scanning 138.67 0 0 0
Characterisation
/NEOs
84.74 0 0 0
0
50
100
150
200
250
Power(W)
Operational Phase
Scanning
CharacterizaLon/NEOs
Target AcquisiLon
Orbital maintenance
CommunicaLons

Solar Arrays Sizing Georgios Galanos
36
Target Acquisition 0 14.23 61.00 61
Orbit maintenance 0 22.62 99.94 96.94
Communicaitons 0 207.94 149.24 148.53
Total 223.41 244.9
Table 6-10 Battery Specifications for operational phase
The study calculates the maximum power the battery can supply in the course
of the operational phase. By using the batteries to split the Power requirements
on the operational phase the area of the solar arrays decreases by 0,3m2
and
the cost by 7000$. This is an addition reduction to the area and cost of the
system after using the battery for peak power demands. The total reduction can
be found in Figure 6-1 Solar arrays comparison
7 Solar Arrays Sizing
Nine basic parameters have to be taken into account for the solar arrays sizing:
1. Operational scheme of the mission: This parameter indicates the time when
the solar arrays must provide power to the load during the mission.
• During launch the spacecraft is not required to provide power to
the loads.
• There will be no power supply in the event that both solar arrays
fail.
2. Solar radiation variation during the mission: As the spacecraft orbits the L1
point, the distance between the spacecraft and the Sun will be between
0,9928 and 0,9891AU and the solar radiation will range between 1400,84–
1411,27W/m2
. This design is using the minimum value. The study also takes
into account the worst-case scenario, when the distance is almost 1AU and
the solar radiation is 1368 W/m2
during the start-up mode.
3. Incidence angle between Sun and solar arrays: The solar arrays have been
designed to be deployable and able to track the sun at any time of the
mission. To ensure reliability, this study also considers the scenario of facing
an undesired angle. According to the literature (Wertz et al, 2015) most of the

37
missions have ran on a 23.5 degrees value, which is the chosen angle for the
design of the solar arrays for the RIICS project.
4. Cells conversion efficiency: The efficiency of the silicon technology is
estimated to be 14,8% (Wertz et al, 2015).
5. Performance degradation: The degradation of the Silicon technology cells is
estimated to be 3,75% per year (Wertz et al, 2015).
6. Inherent degradation: The solar cells are located on a substrate, which
usually results in a 0,77 factor of losses of the solar arrays substrate area
(Wertz et al, 2015).
7. Losses due to transmission inefficiencies: For a Direct Transfer System the
losses due to transmission inefficiencies are projected to be 20%. This
amount of losses is thermally dissipated in the distribution process (Wertz et
al, 2015).
8. Power consumption of each subsystem and component: The power variation
of the solar panels that needs to be provided in the course of the mission
starts from 50.5W and can get up to 222,.W.
9. Power to be provided by the battery: The amount of the power that batteries
need to provide to the system ranges from 4.01W to 171.5W.
After examining all the parameters (the design selection, design selection
analysis and worst-case scenario) the study proceeds in the calculation of
the final size of the solar arrays. In the worst-case scenario, the solar arrays
must be able to provide 222.1W of power to the loads. This leads to a total
3.07m2
area of solar panels and 125,930$ cost. The main features of the
solar arrays are displayed on the table below
Solar Arrays Features
Solar cell efficiency 0.148
DET system efficiency 0.8
Inerent degradation 0.77
Solar cell degradation per year 0.0375
Sun angle (deg) 23.5
Specific cost ($/W) 378

38
Maximum Average Power (W) 222.1
Mass (kg) 7.05
Area (m^2) 3.07
Cost ($) 125,930
Table 7-1 Solar arrays main features
7.1 Battery’s Component Selection
The battery component was selected from the available space applications
market, so to achieve the minimisation of the cost. However, the cost of the
batteries is not available in this study. The calculations were made according to
Patel’s references. A 0.3% per day of self-discharge has also been taken into
consideration (Broussely Pistoia and Knovel, 2007).
The trade-off was made between the products that are provided by SAFT. VES
100, VES 140 and VES 180 were included in the trade-off (Saft, 2008).
To select the appropriate component, the study calculates the total cycles of the
battery to ensure that the selection will be made by the most efficient approach.
The component must provide the necessary energy to the closest approach in
order to reduce the required quantities regarding the cycles of the mission. The
total cycle life of the battery must be over 1881 cycles. This amount of cycles of
life is adequate for all operations from BOL to EOL, however it excludes
emergencies. Case of emergencies will be included later on.
The PCDU will regulate the 28V bus and examine the effects of having lower
battery voltage on the operation of the mission. Given that no price for the
battery has been provided, VES 140 (Saft, 2008) is the selected component as
it is matching the appropriate requirements of the mission. Appendix 9C.2.3
shows the technical specifications of the cells.
7.1.1 Configuration
The specific energy of the VES 140 is sufficient for one cell to provide the
required energy to the spacecraft. As a result, the topology of the battery will be
designed in series so to reach the desired voltage.

39
The nominal voltage of each cell connected in series determines the total output
voltage. “A general guideline is to place cells in series to make the nominal
battery voltage during discharge equal to 80% and during charge about 93% of
the bus voltage” (Patel, 2005), meaning that to reach the 28V requirements 7
cells need to be placed into series. Having less or more than 7 cells will impact
the total mass of the battery and area of the solar arrays as the battery will
require more energy to charge.
The final selection of the cells’ number in series is made by taking into account
the possibility of failure of one cell. The failure of a cell will drop the output
voltage from 25.2V to 21.6V. The failure of one cell will allow the normal
operation of the battery, which means that the system will still be able to provide
the required power to the loads, given that the battery discharge and charge
regulators have approximately 21.6V output (Chapter 8).
The battery is typically designed based on the standards of the worst-case
energy demand scenario during the mission. In the RIICS project the battery is
designed to provide the required peak power in any phase and necessary
power when both solar arrays have failed (worst-case emergency). The battery
will then be charged on Sun safe Mode for 6,7hours. The table below indicates
all the parameters to design the battery according to worst-case scenarios.
Battery's Worst Case Scenario Parameters
Total Battery Discharge 791.16
Total Power Needed to charge the battery in 6,7 hours 171.51
Capacity needed 988.95
Power available during Sun Safe Mode 171.60
Table 7-2 Battery’s worst Case Scenario Parameters
The 7 series cell with the VES 140 component can provide 996.66Whr of
energy, 8Whr higher than the required without the need of parallel configuration.
For the above reasons the topology of the battery pack will be 7 cells in series.
As mentioned in previous chapters, the total battery life must be more that 1881
cycles. Lithium batteries have an average of 1500 cycles at 80% Depth of

Power Control and Distribution Unit (PCDU) Georgios Galanos
40
Discharge (Battery University, 2019). That means that the spacecraft must
accommodate two secondary batteries and an extra battery for redundancy.
The total quantity is three battery packs. The basic specifications of the VES
140 are the following:
Battery VES 140
Specific energy
(Wh/kg)
126
Mass per module (kg) 1.13
Energy (Wh) 142.38
Capacity (Ah) 39
Discharge voltage (V) 3.6
Charge Voltage (V) 4.1
Cells in series 7
Cells in parallel 1
Total capacity (Ah) 39
Total energy (Wh) 996.66
Total mass (kg) 7.91
Table 7-3 VES 140 Shaft Battery
As the power system includes 3 packs of batteries the total mass will be
23.73kg. The cost for each pack of battery is estimated to be around 2,300$
(Patel, 2005)
Finally, since the available power and capacity of the battery are higher than the
required, the batteries will not be installed with oversized margins. Therefore,
the design maintains a cost low.
8 Power Control and Distribution Unit (PCDU)
A PCDU must be able to control and distribute the power through the mission
according to the spacecraft’s needs. For that reason, a PCDU contains a
number of modules, which assist the controlled power distribution. Generally, it
is more convenient to select a PCDU from the existing market. The PCDUs
available in the market are not suitable for the power system, therefore this
project builds the PCDU from components provided by TERMA.

41
To build a proper PCDU for the mission, TERMA provides a Modular Medium
Power Unit concept, which is designed for observation, navigation, science or
low power spacecrafts. The modules used to build the PCDU are all plugged in
a backplane motherboard and can be removed or replaced without any internal
wiring. The module has been designed for a 28V regulated bus with one-single
failure tolerant. Finally, the PCDU can accommodate 21 modules (Terma
space, 2012d).
The components of the PCDU are:
1) 3 Shunt Regulation Modules (S4R) (Terma space, 2012f)
2) 2 Battery Charge / Discharge Regulator Modules (Terma space, 2012a)
3) 3 Equipment Power Distribution Modules (Terma space, 2012b)
4) 4 Heater Power Distribution Modules (Terma space, 2012c)
5) 2 Pyro Firing Drive Modules (Terma space, 2012e)
6) 2 Command and Monitoring Modules
7) 1 Backplane
1. Shunt Regulation Module (S4R): During the mission the available power
from the solar arrays often differs. In many occasions the power that each
loads requires varies. For that reason it is critical to include a regulator to
the system in order to control the available power and protect the loads and
the bus by switching in and out segments of the solar arrays. The individual
segment in the shunt regulator module can be grounded to achieve the
switching out. The configuration of the S4R is a sequential switching shunt
switch regulator module and it accommodates four independent shunt cells.
This approach has the ability to feed the main bus section current via a
series diodes by a parallel switch or to feed the battery section current via
series switch and two series diodes (Terma space, 2012f). This approach is
ideal for the RIICS spacecraft as the dissipated power is directly feeding the
battery. The S4R module provided by TERMA is being designed to provide
an output power capability of 600W (more than the required output power
for the RIICS mission). Missions that have used the TERMA S4R:

42
• Galileo IOV
2. Battery Charge/Discharge Regulator Module (BCDR): The module provided
by TERMA includes two power regulators, a battery Charge Regulator
(BCR) and a battery Discharge Regulator (BCDR). The system is designed
to target a Lithium battery configuration. This module consists of two battery
charging and discharging functions. The first function, which is called End
Of Charge (EOC), is the function that sets the charging limits. The battery is
charging until it reaches the maximum selected level of voltage, after this
point the battery is ready to be discharged. It is important to notice that a
telecommand adjusts the EOC voltage to eight different levels. The second
function regulates the discharge power to the main bus until the voltage
reaches the limits that have been set. This function is called End Of
Discharge. The battery at this stage is not allowed to re-enter a discharge
mode until it reaches again the minimum level of the EOD. The output
power capability of this module is 300W (higher than the required). The
Lithium battery system is been designed in a 7s topology. With this method
the nominal output voltage of the battery is 25.2V and 21.6V in the case of
cell failure. Additionally, the BDR module operates in a range of 0–25.6V
and BDR with the range of 16–27V. Therefore, in the case of failure of one
cell, the PCDU will be able to operate normally. The voltage output of the
battery is often lower than the voltage output of the bus in PCDU designs.
The BDR is a conventional step-up regulator (Jensen and Laursen, 2002)
that increases the output voltage. Missions that have used the TERMA
BCDR are the following:
• Mars Express
• Rosetta
• Venus Express
3. Equipment Power Distribution Module (EPD): This module provides a
number of protection switchers. The PCDU must be designed in a way that
all the spacecraft loads are equipped with one switch protection. The
current module, which is provided by TERMA, consists of sixteen Latching

43
Current Limiters (LCL) and can be applied to the loads. There are twenty
loads on the spacecraft and for that reason three Modules are needed of
which one is for redundancy. The function of this module is protecting the
upstream main power bus from an overload or short circuit current.
Missions that have used the TERMA EPD are as follows:
• XMM-Newton
• Integral
4. Heaters Power Distribution Modules (HPD): This module is responsible for
the distribution of the bus power to the spacecraft heaters. At the moment
TERMA provides this module with sixteen output switches divided in two
groups of eight switches. Upstream LCL protects each of these groups. The
RIICS spacecraft includes thirty six heaters, which leads to the selection of
four HPDM of which one is for redundancy. Missions that have used the
TERMA HPD are the following:
• XMM-Newton
• Integral
5. Pyro Firing Drive Modules (PFD): This module is responsible for turning on
and off the valves of the thrusters. One extra module is for redundancy.
Missions that have used the TERMA PFD:
• Hispasat
6. Command and Monitoring Modules (CM): This module introduces the
communication interface to the data management subsystem. Two modules
are being accommodated, the one for redundancy. Missions that have used
the TERMA CM are:
• Mars Express
• Rosetta
• Venus Express
7. Backplane: This module interconnects the aforementioned modules and
closes the PCDU.

Conclusion Georgios Galanos
44
The main final specifications of the PCDU are the mass, width, height and
length which are 11.37kg, 235mm, 156mm and 379mm respectively.
9 Conclusion
The following conclusion summarises the final design of the Electrical Power
system.
The distribution of the power demand to the loads is achieved with a fully
regulated 28V bus working with a DET system.
Two deployable 3.07m2
solar arrays are placed on two rigid structures and
support the spacecraft with the required power during the mission. This
excludes the worst-case scenario of emergency (discussed in previous
Chapter). The two solar arrays cost 125,930$ and weigh 7.07kg.
During peak power, scanning and characterisation/detection NEOs modes and
emergencies, a battery supplies the necessary power demand. Three units of
batteries are projected to cover the entire mission, of which one battery is for
redundancy. The battery consists of a 7s cell configuration and weighs 7.91kg.
The total cost is around 6,900$ and total mass is 23.73kg.
A power control and distribution unit is being attached to ensure the efficient
operation of the solar arrays and battery and the successful and safe power
distribution to the loads. The PCDU weighs 11.37kg.
The total approximate cost of the Electrical Power will be around 133,000$ and
total weigh will be 42.17kg.
The Electrical Power system successfully covers the corresponding
requirements set by the project’s system engineers.
Further analysis of the solar cells and their configuration is essential. Studying
the sensitivity of the battery regarding temperature, actual voltage requirements
for each subsystem and instrument of the mission and more possible efficient
designs of the electrical power system is desired.

REFERENCES Georgios Galanos
45
REFERENCES
Angrist, S. W. (1982) Direct Energy Conversion, 4th edn, Copyright Allyn and Bacon,
New York)
Antonio De Luca (2011). Architectural Design Criteria for Spacecraft Solar Arrays,
Solar Cells - Thin-Film Technologies, Prof. Leonid A. Kosyachenko (Ed.), ISBN: 978-
953-307-570-9, InTech, Available from: http://www.intechopen.com/books/solar-cells-
thin-film-technologies/architectural-design-criteria-for-spacecraft- solar-arrays
Battery University (2019) Comparison Table of Secondary batteries available at:
https://batteryuniversity.com/learn/article/secondary_-batteries (accessed February
2019)
Ben Johnson (2012) Power Sources for Space Exploration, Stanford University
available at: http://large.stanford.edu/courses/2012/ph240/-johnson1/ (accessed in
November 2018)
Broussely, M., Pistoia, G. and Knovel, (2007), Industrial applications of batteries, 1st
ed., Elsevier, Amsterdam ; Boston.
Clyde Space (2010), Secondary Batteries, available at:
http://www.clydespace.com/resources/powerschool/power_storage/secondary_batterie
s, (accessed January 2019).
David W. Miller and John Keesee, Spacecraft Power System, MIT OpenCOurseWare
available at: https://ocw.mit.edu/courses/-aeronautics-and-astronautics/16-851-
satellite-engineering-fall-2003-/lecture-notes/l3_scpowersys_dm_done2.pdf, (2003)
(accessed December 2019)
Fortescue, P. W., Stark, J. and Swinerd, G. (2003), Spacecraft systems engineering,
3rd ed, Wiley, Chichester.
Huynh, P. T. and Cho, B. O. H. (1999), "Design and analysis of a regulated peak-power
tracking system", IEEE Transactions on Aerospace and Electronic Systems, vol. 35,
no. 1, pp. 84-92.
James R. Wertz, David F. Everett and Jeffery J. Puschell (2015), Space Mission
Engineering: The New SMAD, Microcosm Press, USA; Hawthorne.
Jensen, H. and Laursen, J. (2002), "Power conditioning unit for Rosetta/Mars express",

REFERENCES Georgios Galanos
46
6th European Space Power Conference, 6 May 2002 through 10 May 2002, Porto, pp.
249.
Jiang, J. -., Huang, T. -., Hsiao, Y. -. and Chen, C. -. (2005), "Maximum power tracking
for photovoltaic power systems", Tamkang Journal of Science and Engineering, vol. 8,
no. 2, pp. 147-153.
Jurian, T. 2019 RIICS: Rapid Imminent Impactor Characterisation System -
Configuration & Structures
Landis, G. A. (2006) Reevaluating Solar Power Systems for Earth, IEEE 4th World
Conference on Photovoltaic Energy Conversion 2006 , NTRS-2007-0005136.
Liébana Moradillo, O., 2019 RIICS: Rapid Imminent Impactor Characterisation System
- Operations and Mechanisms
Niels E. Jensenh (2003), Satellite Power System, ESA available at:
http://www.esa.int/esapub/br/br202/br202.pdf (accessed December 2019)
Patel, M. R. (2005), Spacecraft power systems, CRC Press, Boca Raton.
Picavez, C. (2019) RIICS: Rapid Imminent Impactor Characterisation System – Launch
System
Saft (2008), VES 180 - Rechargeable lithium battery datasheet, available at:
http://www.saftbatteries.com/doc/Documents/space/Cube712/VES%20180.e9cf5d8f-
3cbd-4921-8ac0-89d7b13bd0c0.pdf (accessed March 2019).
Salina Asif (2008) Evolutionary computation based multi-OBJECTIVE design search
and optimization of spacecraft electrical power subsystem, University of Glasgow,
available at: http://theses.gla.ac.uk/373/1/2008AsifPhD.pdf
Seurin, N. 2019 RIICS: Rapid Imminent Impactor Characterisation System - System
Engineering Requirements & Risks
SPS Concept Development and Evaluation Programme Reference System Report
(1978) US DOE and NASA DOE/ER 0023.
Terma space (2012a) Battery C/D Regulation Module, available at:
https://www.terma.com/media/177689/battery_cd_regulation_module.pdf (accessed
February 2019)
Terma space (2012b) Equipment Power Distribution Module, available at:

BIBLIOGRAPHY Georgios Galanos
47
https://www.terma.com/media/177695/equipment_power_-distribution_module.pdf
(accessed February 2019)
Terma space (2012c) Heater Power Distribution Module, available at:
https://www.terma.com/media/177698/heater_power_distribution_module.pdf
(accessed February 2019)
Terma space (2012d) Modular Medium Power Unit, available at:
https://www.terma.com/media/150039/modular_medium_power_unit.pdf (accessed
February 2019)
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https://www.terma.com/media/177719/pyro_firing_drive_module.pdf (accessed
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BIBLIOGRAPHY
Bewick, R., CUTE, Communication and Power Subsystems. Group Design Project
report for MSc in Astronautics and Space Engineering, Cranfield University, 2009.
Busquets Corominas J., Marco Polo ‘Lite’, Systems engineering (Requirements,
baseline and coordination) and Electrical power subsystem. Group Design Project
report for MSc in Astronautics and Space Engineering, Cranfield University, 2010.
Lim, Timothy M., (2016), "A modular electrical power system architecture for small
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Zane Brough, Claudio Paoloni, (2015), Advanced Deployable/Rectractable Solar Panel
System for Satellite Applications, World Academy of Science, Engineering and
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48
6b70fa7f8fa.pdf?_ga=2.166006782.1760230267.15435815401910833425.154358154
0

Executive Summary: RIICS - Electrical Power Subsystem Georgios Galanos
49
APPENDICES
Appendix A Executive Summary: RIICS - Electrical
Power Subsystem
The Appendix A summarise the Electrical Power subsystem work package.
The main features of the Electrical Power subsystem are:
1. Two Silicon rigid deployable solar arrays, each consist of three panels.
To be able to provide the power requirement (50.5W – 222.1W) the area
of each of the array must be 1.535m2
.
2. Three pack of battery (one for redundancy). The total capacity of the
battery is 998Wh and is consists of 7 cells in series provided by saft
(VES 140 model). The total weight of the batteries is 23.73kg.
3. The total approximate cost of the Electrical Power will be 133,000$ and
weights 42.17kg.
4. A PCDU will control and distribute the power generated and will weight
11.37kg.
5. A fully regulated bus operating up to 28V.
6. No eclipses during the mission.
7. 6 years designed lifetime.
The electrical power subsystem was designed to be able to provide the required
power during the whole mission of the spacecraft in the most efficient and low
cost approach. The cost of each of the components may be different in reality.
References have been used for calculating the approximate cost of the
electrical power system and the values are not reliable.

Common Appendix Georgios Galanos
50
Appendix B Common Appendix

B.1 Mission objectives
Mission statement
“The aim of the mission is to build, in partnership with Aistech Space, a Physical
Characterization and Scanning System for small Rapid Imminent Impactors”
B.1.1 Primary objectives
● Characterization of imminent impactors to improve the understanding of the
small Near-Earth Objects by comparing information obtained pre-impact from
the spacecraft and post-impact information
● Detection of imminent impactors that are not detectable from the ground by
using the spacecraft as an orbital scanning station
B.1.2 Secondary objectives
● Detection of Residual Space Objects (RSOs) and utilisation of Space Surveillance
Network (SSN) database
● Classification of NEOs and detection of a range of compounds on NEOs surface
● Exoplanets characterisation through photometry and spectroscopy
B.2 Key mission requirements
B.2.1 Mission performance
● Mission_01 - One third of all NEOs not detected by Earth with a closest
approach distance to Earth <= 0.03 AU with absolute magnitude H <= 30
shall be characterized
● Mission_02 - The s/c shall be able to obtain the light curve of the target
● Mission_03 - The s/c shall be able to obtain the reflectance spectrum of the
target
● Mission_04 - The s/c shall be able to observe the target in the Near Infrared
(0.7 μm < 𝜆 < 2.5 μm) and Visible spectral regions (0.4 μm < 𝜆 <0.7 μm)

51
● Mission_05 - The s/c shall be able to observe targets not detected by
terrestrial telescopes
B.2.2 Mission constraints
● Political_Constraint - The mission shall satisfy a hypothetical mission call
from ESA’s Space Situational Awareness (SSA) Program
● Cost_Constraint - The total RIIC cost should be 50 M€
● Schedule_Constraint - Small-Class mission: the mission should be
implemented under a fast scheme of 5 years
● Development_Constraint - The Technical Readiness Level (TRL) should be 5-
6 (ISO scale) by the end of the short preparation phase and before the
mission adoption
B.3 Risk

● Regarding the mission, the most critical work packages are: Payload,
Communications & OBDH, AOCS and Propulsion.
Highest Criticality Events Key prevention actions
Payload: detector
degradation
-Payload instruments shall be protected against impacts
-The s/c shielding materials should have high hardness
-Add little cameras on the s/c to detect failures
Comms and OBDH: Loss of
scientific and housekeeping
data
-Provide redundancy
-Use dispatched ground stations to avoid interferences
-Mount antennas in less risky part (potential impacts,
radiations…) considering the pointing requirements
-Provide safe-mode in case of anomaly
AOCS: Wrong attitude of
the spacecraft
-Correlate computations with the one of similar past
missions
-Organize frequent team meetings (several per week) to
improve communication and transmit information and
results

52
- Apply margins when designing the propellant budget
-Reduce orbit eccentricity and pointing errors to acceptable
limit
Propulsion: Loss of
spacecraft
-Check the quality of the system
-Add cameras on the s/c to provide follow up of the orbit
phase
-Ensure electrical continuity of the surfaces
-Prevent propellant mixing between opposite tanks
Table B-1 Highest Criticality Events and Key Prevention Actions
B.4 Trade-off
Weigh Geo EW Geo CH L1 EW L1 CH DRO EW DRO CH
Performance 5 1 2 3 5 1 1
Impact 4 5 3 5 3 4 2
Cost 4 4 5 2 4 1 2
AOCS demands 3 3 3 5 5 5 5
Feasibility 2 5 3 3 3 3 3
Environmental
effects
2 3 3 5 5 4 4
Communications 1 5 5 2 3 1 2
Total 105 71 68 76 87 55 52
Table B-2 Trade off

53
B.5 Budgets
B.5.1 Mass Budgets
Table B-3 Mass Budget Breakdown
Component Units Unit mass (kg) Total dry mass
Propulsion module
Primary thruster 1 5.4 5.4
Fuel tank 4 6.81 27.24
Oxidiser tank 2 9.29 18.58
2 Subsystem Mass (Kg)
Payload 80
Power 10.06
Structure 102.36
Mechanisms 34.71
AOCS 55.13
Communications 31.8
Thermal 7.59
OBDH 12.8
Margin 5%
Total dry mass 350
Wet Mass 520.17

54
Pressurant tank 1 6.04 6.04
Pyro valve 9 0.16 1.44
Non return valve 4 0.3 1.2
Fill/drain valve 19 0.09 1.71
Pressure regulator 2 1.72 3.44
Isolation valve 6 0.545 3.27
Filter 7 0.114 0.798
Pressure transducer 10 0.14 1.4
Piping 1 15 15
Structure 1 140 140
Pressurent 1 1.27 1.27
Total dry mass 222
Reaction Control System (RCS)
RCS thruster 16 0.33 5.28
Fuel tank 1 2.66 2.66
Pressurant tank 1 1.98 1.98
Pyro valve 4 0.16 0.64
Non return valve 2 0.3 0.6
Fill/drain valve 7 0.09 0.63
Pressure regulator 2 1.72 3.44

55
Isolation valve 6 0.545 3.27
Filter 2 0.114 0.228
Pressure transducer 6 0.14 0.84
Piping 1 10 10
Pressurant 1 0.03 0.03
Total dry mass 29.60
Attitude Control and Determination System (ACDS)
Star tracker camera 2 1 2
Star tracker processor 2 1.2 2.4
Star tracker baffle 2 0.53 1.06
Sun sensor 2 0.215 0.43
MIMU 2 4.44 8.88
Reaction wheels 4 4.1 16.4
Total dry mass 31.17
Table B-4 AOCS Mass Breakdown
B.5.2 Link Budgets
Parameter Value Unit Comments
Frequency 8.45 GHz X-Band downlink
Transmit Power 100 W X-TWTA
Transmit Power 50 dBm
Transmitter Diameter 1.2 m

56
Peak Transmitter Gain 37.9 dBi
EIRP 85.9 dBm DSN 34m
Propagation path
length
1.748E+06 Km Maximum distance in the orbit from
Earth
Space loss 236 dB
Atmospheric
Attenuation
1 dB
Other losses 1 dB Cables, switches etc.
Receiver Diameter 34 m
Receiver gain 68.9 dBw
System Noise
Temperature
28 K
Eb/No 13.5 dB
Link Margin 3 dB
Required Eb/No 10.5 dB BPSK modulation
Data rate 9.14 Mbps
Table B-5 HGA Link Budget
Parameter Value Unit Comments
Frequency 8 GHz X-Band downlink
Transmit Power 100 W X-TWTA
Transmit Power 50 dBm
Transmitter Diameter 0.5 M
Peak Transmitter Gain 29.8412 dBi

57
EIRP 77.8 dBm ESA 15 m
Propagation path
length
1.74E+06 Km Maximum distance in the orbit from Earth
Space loss 236 dB
Atmospheric
Attenuation
1 dB
Other losses 1 dB Cables, switches etc.
Receiver Diameter 15 m
Receiver gain 50 dBw
System Noise
Temperature
133 K
Eb/No 13.5 dB
Link Margin 3 dB
Required Eb/No 10.5 dB BPSK modulation
Data rate 611.863 Kbps
Table B-6 LGA Link Budget
B.5.3 Power Budget
Phase Subsystem Average
Power (W)
Peak
Power
(W)
Charge Discharge Energy
Discharged
(Whr)
Power
to charge
(W)
Total
Power
(W)
Scanning Payload 56 OFF ON 222,1
Structure /
Mechanics
18
Thermal 44
Comms 13,5
OBDH 33
ADCS 57,6 69 136,67
Propulsion 0
Characterisation / Payload 56 OFF ON 222,1

58
NEOs
Structure /
Mechanics
18
Thermal 44
Comms 13,5
OBDH 33
ADCS 57,6 69 87,74
Propulsion 0
Target
Acquisition
Payload 0 ON OFF 49,6 222,1
Structure /
Mechanics
18
Thermal 44
Comms 13,5
OBDH 28
ADCS 69
Propulsion 0
Orbit
Maintenance
Payload 0 ON OFF 85,54 219,1
Structure /
Mechanics
18
Thermal 0
Comms 13,5
OBDH 28
ADCS 19,5 39
Propulsion 54,56
Communications Payload 0 ON OFF 137,126 221,38
Structure /
Mechanics
18
Thermal 0
Comms 30,01
OBDH 33
ADCS 3,25 39
Propulsion 0
Transfer Payload 0 ON ON 221,8 51,37 145,32
Structure /
Mechanics
18
Thermal 0
Comms 2,5 30,05
OBDH 28

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