Small Satellite Fly-By Mission to the Moon LASAR-SAT Mission PDR
NANOSTAR consortium Project funded by the Interreg Sudoe Programme through the European Regional Development Fund (ERDF)
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NANOSTAR STUDENT CHALLENGE Small Satellite Fly-By Mission to the Moon LASAR-SAT Mission PDR
May 2019, Madrid
Universidad Carlos III de Madrid Master’s Degree in Aeronautical Engineering Space Systems Design.
“We would like to give special thanks to all those Americans who built the spacecraft; who did the construction, design, the tests, and put their hearts and all their abilities into those craft. To those people tonight, we give a special thank you, and to all the other people that are listening and watching tonight, God bless you. Good night from Apollo 11”
Neil Armstrong, July 23, 1963
EXECUTIVE ABSTRACT
This document gathers the Preliminary Design of a nanosatellite fly-by mission to the Moon as a response to the NANOSTAR Student Challenge 2019, in which 7 universities, 2 aerospace clusters and 3 ESA Business Incubation Centres in France, Spain and Portugal are involved.
The main goal of the presented design is to provide a suitable Space Segment which enables the accomplishment of the target mission. This one, pursues the acquisition of science data during the periselenium pass from altitudes above the Moon’s surface as low as 100 km.
Due to the nature of the mission, high accurate systems are to be designed in order to provide an adequate support for the optical camera system payload which will be operative taking images of lunar soil just during the periselenium pass. Therefore, margins on design and redundancy of the critical components have been considered and included in all pertinent computations. In fact, those two aspects are of great relevancy on the performed design. They are considered to aid to minimise the associated risk of the mission, mitigating the undeniable inability to access the space environment and carry out repairs or upgrades.
The performed work is based on a collaborative and flexible but organized methodology, in which own developed tools are used to enhance the parametric analysis and cover of a wider spectrum of design possibilities. Engineering tools such as MATLAB, Microsoft Excel, IDM tools or ESATAN-TMC are considered key ones to conduct the necessary analyses and results monitoring.
The final preliminary design solution comes across after performing several trade-off analysis and iterations on each of the subsystem, evaluating alternative proposals and both updating and verifying low level requirements frequently. The targets of design weighted on the trade-off analysis until a suitable compromise solutions is portrayed, have been mainly focused on balancing factors such as functionality, mass budget, power budget and thermal behaviour, among others.
Therefore, the one proposed by LASAR-SAT engineer’s team encompasses the following mission Quick Facts, which are the result of a design philosophy based mainly on simplicity and reliability.
- Ground Station: Kourou S-Band Station at French Guiana in South America. - Structure characteristics: Overall mass (margins included) of 8.84 kg. Total volume 3U. Van Allen charged particles shielding. - Thruster: Ion-thruster BUSEK BIT-3. - Propellant: Solid Iodine. - Orbit: Apogee raising transfer until Earth-Moon distance. - Mission duration: 18.4 months 30 h in close Moon phase. Positive fly-by on the order of minutes. - Power: Daylight: 54.5W Eclipse: 14.75 W Solar arrays 0.123m2 GaAs Lithium-Ion Batteries - AOCS 3-axis stabilization Sensors: Startracker + Gyro, Sun sensors Actuators: RW + desaturation system - Thermal Control Passive approach.
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TABLE OF CONTENTS
EXECUTIVE ABSTRACT ...... I List of Figures...... IV List of Tables ...... V ACRONYMS ...... VI 1. INTRODUCTION ...... 1 Team organisation ...... 2 Work Logic ...... 2 2. MISSION OVERVIEW AND REQUIREMENTS FLOWDOWN ...... 3 State of the Art ...... 3 System segments and mission goals ...... 4 Requirements Flowdown ...... 5 3. SUBSYSTEMS ANALYSIS AND DESIGN ...... 6 3.1 Mission analysis ...... 6 Delta-V budget ...... 6 Spacecraft trajectory ...... 8 3.2 Systems operations modes ...... 12 3.3 Space propulsion Subsystem ...... 12 Chemical Vs. electrical propulsion ...... 12 Refinement process of the electrical thruster ...... 15 3.4 Attitude, determination and control subsystem ...... 18 Objectives of the ADCS ...... 18 Disturbance Torque Environment ...... 18 Design Solution for the ACDS ...... 20 Sensor Selection ...... 26 Control Algorithm ...... 28 3.5 Communication Subsystem and Ground Segment ...... 29 Summary of the main S/S constraints/ Assumptions ...... 29 Communications configuration selection ...... 29 Ground Station selection ...... 31 Antenna Selection ...... 31 Link Budget Calculation and trade-off ...... 31 3.6 Electric Power subsystem ...... 33 Summary of the main S/S constraints ...... 33 Selection of the EPS configuration ...... 34 Summary of the selection ...... 36 Secondary Power Source Sizing ...... 36 3.7 Mechanical design and structure...... 37 Launcher selection and launch phase description ...... 37 Configuration trade-off ...... 38 Mission Phases ConfigurationS ...... 41 Mass budget ...... 41 Center of mass and Inertia Tensor ...... 42 Natural Frequencies ...... 42 Mechanisms ...... 43 Structural Materials ...... 44 3.8 Thermal control subsystem ...... 45 4. RISK ANALYSIS AND MITIGATION ...... 49 5. CONCLUSIONS AND FUTURE WORK GUIDELINES ...... 50
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6. REFERENCES ...... 51 APPENDICES ...... 53 Appendix A – Report Change Record Table ...... 53 Appendix B – Requirements ...... 54 Subsystem Requirements Change Record Table ...... 54 Communications S/S ...... 55 Electric Power S/S ...... 56 Space Propulsion S/S ...... 57 Attitude Detremination and Control Determination S/S ...... 58 Structure, Mechanical and Mechanisims S/S ...... 59 Thermal Control S/S Requirements ...... 60 Appendix C – Minutes of Meetings, MoM’s...... 61 Appendix D – Trade-Off Studies ...... 66 Appendix E – Risk Analysis/ Mitigation Table ...... 68 Appendix F – Launch Phase Description ...... 71 Appendix G – Design and Structire Details ...... 74 Appendix H – Configurations mass Budget ...... 78 Appendix I – Additional Telecommunications s/s information ...... 81 Appendix J – Additional information of the Thermal S/S ...... 83
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LIST OF FIGURES
Figure 1 - OBS ...... 2 Figure 2 - Space System Segments ...... 4 Figure 3 – Delta-V Budget ...... 8 Figure 4 - Transfer Trajectory ...... 10 Figure 5 - Fly-By Trajectory ...... 10 Figure 6- Fly-By Velocity Vectors ...... 11 Figure 7 – Disposal Trajectory ...... 11 Figure 8 - Propellant Mass Budget ...... 14 Figure 9- Fly-By Trajectory Definition ...... 22 Figure 10-Control Algorithm Schematic SMART-1 ...... 28 Figure 11 - PDI algorithm ...... 28 Figure 12 - Roll Error ...... 28 Figure 13-QPSK Modulation Symbols ...... 30 Figure 14 - Atmospheric Losses ...... 30 Figure 15 - Power Distribution along Transfer Orbit ...... 34 Figure 16 - Primary Power Source Selection ...... 34 Figure 17 - ISIS 6-Unit DuoPack ...... 37 Figure 18 - First Configuration Views ...... 38 Figure 19 - Second Configuration...... 39 Figure 20 - Third Configuration (Science Mode) ...... 39 Figure 21 – Subsystems/Components Distribution ...... 40 Figure 22 - Transfer Mode ...... 41 Figure 23 - Launch Mode ...... 41 Figure 24 - Detail of XATCOBEO SA Deployment Mechanism [17] ...... 43 Figure 25 - Detail of Camera Riel System Articulation...... 44 Figure 26 - Components Temperature Evolution for the Eclipse (Left) and Sun Light (Right) Scenarios ...... 48 Figure 27 - Schematic Soyuz Launch [15]...... 71 Figure 28 - Timeline of Soyuz Launch Sequence [15] ...... 71 Figure 29: Schematic Gto- reaching.[15] ...... 72 Figure 30- Typical Sequence of events during Payload deployment Stabilization ...... 73 Figure 31 - Propulsion Subsystem ...... 74 Figure 32 - Power Subsystem ...... 74 Figure 33 - ADCS Subsystem ...... 75 Figure 34 - OBC ...... 75 Figure 35 - Communications Subsystem ...... 76 Figure 36 - Structure Subsystem ...... 76 Figure 37- Kourou Station ...... 81 Figure 38 – selected Antenna Gain Evolution ...... 82 Figure 38 - Orbital Results for the Eclipse (At Perigee) Scenario ...... 83 Figure 39 - Orbital Results for the Sun Light Scenario ...... 83 Figure 40 - Orbital Results for the Hypothetic Scenario With Eclipse At Apogee ...... 84 Figure 41 - Nodes Temperature Evolution for the Eclipse (At Perigee) Scenario ...... 84 Figure 42 - Nodes Temperature Evolution for the Eclipse at Apogee Scenario ...... 85 Figure 43 - Nodes Temperature Evolution for the Sun Light Scenario ...... 85
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LIST OF TABLES
Table 1 - Mission Requirements (High Level) ...... 5 Table 2- FEEP Thrusters Trade-Off for Wet Mass of 9kg ...... 15 Table 3 - Bit-3 Results for 9kg Wet Mass ...... 17 Table 4 - Propellant Mass Budget and Propulsion data ...... 17 Table 5- Design Drivers ...... 23 Table 6 – Reaction Wheels Catalogue ...... 24 Table 7 - RCS Evaluated Systems Information and Results ...... 25 Table 8- Final Solution for Desaturation ...... 26 Table 9 - Sensor Characteristics ...... 27 Table 10- Selected Sensors and Back-up system ...... 28 Table 11- Link Budget ...... 32 Table 12-Antenna Specifications ...... 32 Table 13-transceiver specifications ...... 32 Table 14 - Power Requirements ...... 33 Table 15 - Power Requirement during Daylight and Eclipse ...... 34 Table 16 - Daylight and Eclipse Length ...... 35 Table 17 - Power available for both configurations at BOL & EOL ...... 35 Table 18 - Solar Panel and Equipment Specifications ...... 36 Table 19 - Battery Specification ...... 36 Table 20 - ISIS 6-Unit DuoPack operational characteristics ...... 37 Table 21 – Components Flow-down ...... 40 Table 22 - Launch Configuration Mass Budget ...... 42 Table 23 - Launch Configuration Cg & Inertia ...... 42 Table 24 - Transfer Configuration Cg & Inertia ...... 42 Table 25 - Science Configuration Cg & Inertia ...... 42 Table 26: Data for Frequencies Calculations ...... 43 Table 27: Appendages Natural Frequencies ...... 43 Table 28- Structural Materials Properties ...... 44 Table 29 – Heat Dissipation Definition ...... 46 Table 30 – Thermal Materials Definition ...... 46 Table 31 – Coating Properties ...... 47 Table 32 – OFF-ON Status at both scenarios ...... 47 Table 33 - Probability Impact Risk Scoring ...... 49 Table 34 – PDR Document Change Record ...... 53 Table 35 - SS Requirement Change Log Record ...... 54 Table 36 - Communications S/S Requirements ...... 56 Table 37 – Electric Power S/S Requirements ...... 57 Table 38 – Space Propulsion S/S Requirements ...... 57 Table 39 – ADCS S/S Requirements ...... 58 Table 40 – Structure, mechanical and mechanisms S/S Requirements...... 59 Table 41 – Thermal Control S/S Requirements ...... 60 Table 42 - MoM 1 ...... 61 Table 43 - MoM 2 ...... 62 Table 44 - MoM 3 ...... 63 Table 45 - MoM 4 ...... 64 Table 46 - MoM 5 ...... 64 Table 47 - MoM 6 ...... 65 Table 48 - Evaluation Criteria for Propulsive Trade-Off Analysis ...... 66 Table 49 - Evaluation Criteria and Results for Telecommunications Trade-Off Analysis ...... 66 Table 50 - Weighted Parameters for ADCS sensors Trade-Off Analysis ...... 67 Table 51 - Weighted Parameters for ADCS Desaturation system Trade-Off Analysis ...... 67 Table 52 - Weighted Parameters for ADCS RW Trade-Off Analysis ...... 67
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Table 53 – Weighted Parameters for Mechanical Trade-Off Analysis ...... 67 Table 54 - Risk Analysis / Mitigation ...... 70 Table 55 - Launch Configuration Mass Budget ...... 78 Table 56 - Transfer Mode Mass Budget ...... 79 Table 57- Science Mode Mass Budget ...... 80
ACRONYMS
AOCS/ADCS Attitude and Orbit Control Subsystem (Attitude Determination Control Subsystem) BOL Beginning of Life CE Concurrent Engineering cg. Center of gravity CPU Computer Processing Unit CFRP Carbon Fiber Reinforced Polymer DOD Depth of Discharge ECI Earth Centered Inertial EOL End of Life EPS Electrical Power Subsystem ESA European Space Agency FEEP Field Emission Electric Propulsion FOV Field of View GS Ground Segment GNSS Global Navigation Satellite System GRMS Gaussian Root Mean Square Vibration GTO Geostationary Transfer Orbit HET Hall Effect Thruster HDRA Hold-Down and Release Actuator. IMU Inertial Measurements Unit IR Infrared ISIS Innovative Solutions in Space ISL Innovative Space Logistics Isp Specific Impulse LEO Low Earth Orbit OBC On Board Computer PDR Preliminary Design Review PPU Power Processing Unit PPT Pulse Plasma Thrusters QAM Quadrature Amplitude Modulation QPSK Quadrature Phase Shift Keying RCS Reaction Control System REACT Resettable non-explosive Actuator RIT Radio-Frequency Ion Thrusters RW Reaction Wheels SA Solar Arrays S/C Spacecraft S/S Subsystem SOI Sphere of Influence SP Solar Pressure TCM Trajectory Correction Maneuvers TTC Telemetry and Telecommand UHF Ultra High Frequency
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1. INTRODUCTION
The aim of the present document is to gather a Preliminary Design Review (PDR) of a nanosatellite space mission to the Moon. The satellite, is to be equipped with a scientific payload, consisting on a camera which will observe and make some measurements of the Moon's surface at a close distance during a fly-by.
Henceforth, the scope of the report is to propose a feasible solution to this challenge, including the definition and sizing of the most relevant subsystems. Those, are defined in a relatively still flexible manner that provides acceptable estimations serving as a starting point for further, more detailed, analysis carried out along the following design phases. That is, the present PDR serves as a conclusion of the project phases A and B, according to ESA’s guidelines [6]. A preliminary mission analysis and performance estimation of the overall system is therefore provided, justifying that the present solution satisfies the top-level mission requirements and deals with the constraints.
In relation to the core of the present document, the sections enumerated at the List of Contents are to be explained to serve as a guide for project understanding:
- Section 1: Includes the document scope and level of detail in the performed design. Within this section, team organisation and consequently, the definition of responsibilities and work philosophy are also gathered. Besides, a brief summary of the following sections is provided. - Section 2: Devoted to the mission overview, explaining its reason to be, main imposed requirements and preliminary concept of the mission as well as the found solution. Moreover, some historical background on previous and similar exploration missions related to the Moon are included since they have been consulted to deeply understand the design goals. - Section 3: Subsystems analysis and preliminary selection of its components based on the flow-down requirements imposed. To support the subsystem sizing, the used theoretical background is presented (referenced to literature on the subject) along with the summaries of the main results. The six assessed subsystems are evaluated for different configurations so the required trade-off analysis are also presented herein. Notice that just the conclusions of such comparisons are provided at the section, not the tables gathering the weighting factors. In particular, the following aspects are treated along the section. . Mission analysis to estimate the Delta-V budget. . Operative modes of the system. . Propellant mas budget along with the selected thruster technology. . ADCS selected set of sensors and actuators. . Ground segment proposal followed by the communications subsystem sizing. . Designed sources of electric power both primary and secondary. . Structural configuration of the space segment to estimate the mass budget, inertias and natural frequencies for distinct mission phases, ensuring suitability of launching phase. Finally, the proposed space segment solution is presented as an IDM-CIC file, which contains the 3D model of the spacecraft with a characterization of all its subsystems in terms of mass and position. . Thermal analysis of the selected configuration for different worst case scenarios to ensure proper performance of the components within their temperature ranges. - Section 4: Devoted to the risk analysis and proposal of tailored response actions. - Section 5: Final conclusions extracted from the PDR development are stated, highlighting the more relevant design aspects and future lines of work. Moreover, non-compliances with low level requirements which need further level of detailed design are presented with possible solutions. - Appendices: Encompasses all additional information used along the design. For instance complete requirements flow-down, MoM’s, launcher selection, GS information, additional 3D images of the configuration, risk analysis, trade-off criteria and intermediate results, etc.
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TEAM ORGANISATION
The team in charge of carrying out the Preliminary Design gathered at the present report is formed by five Graduated in Aerospace Engineering. Four of them are in charge of the main four S/S treated at this early stage of the design: Telecommunication & Power, Propulsion & ADCS, S/C Configuration, Structure & Mechanisms and lastly, Thermal Control. Besides that, the last team member carries out the Project Manager role, being the ultimate responsible of the PDR and acting as a focal reference point to the rest of team constituents.
Based on the background experience of each of the team members, the roles distribution is tailored and the Organisation Breakdown Structure of the team defined as shown at Figure 1, where the different S/S responsibles are identified. Nevertheless, it is relevant to highlight that the design of the particular subsystems cannot be evaluated alone, meaning that although the subsystems isolation has been pursued by means of specific low level requirements, a collaborative philosophy must be put into practice within the design, linking the S/S. Thus, linking the performance of the engineers in charge of those S/S. Please, refer to the Appendices section to consult the particular list of duties that have been carried out by each of the engineers.
FIGURE 1 - OBS
WORK LOGIC
Concurrent engineering has define the main work philosophy employed along the project. By using this technique, there is an early consideration for every aspect of the product development process, saving time, resources and cost. Although the main project responsibility is place on the PM, the truth is that decisions are tried to be taken, the majority of times, in parallel. Therefore, accounting for the acquired knowledge of the five engineers. In fact the multidisciplinary nature of the design makes other type of work philosophies barely suitable
Based on this idea, an early division of work load has been carried out. Different Work Packages associated to each of the subsystems were identified to provide a more controllable and manageable way to monitor the design development, since it is easier to evaluate highlights in reduced batches. Their monitoring has been mainly done by means of meetings, where topics such those found at the Appendix C – Minutes of Meetings, MoM’s. were treated.
Finally, and for the sake of brevity, just a few working and communication tools employed along the design are to me mentioned. Those are: Google Drive, Gmail, MATLAB, Microsoft Office tools (Word, Excel and PowerPoint), ESATAN-TMS and Integrated Design Model (IDM) tools and viewers (developed by CNES).
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2. MISSION OVERVIEW AND REQUIREMENTS FLOWDOWN
One of the first stages of every mission design is the understanding of its reason to be. This way, engineers can easily tailor their designs to mission goals accomplishment from the very beginning. Therefore it is important to highlight why the Moon seems to be so appealing in the last years among the scientific community.
In fact it is due to the discoveries found by the NASA’s Lunar Prospector in 1999 related to certain concentrated hydrogen signatures which were detected in permanently shadowed craters near the lunar poles. The indication of lunar water could have far-reaching implications as humans expand exploration for instance in their will to explore solar system. Nevertheless, since the Moon’s gravity is less than one fifth of Earth’s gravity, the Moon has practically no atmosphere so that any light elements or compounds deposited on the surface are subject to direct exposure to the vacuum of space and solar radiation. That is the reason why the acquisition of data to confirm water existence and determine the level of resources that could be available to human usage, is still under investigation, leading to a large number of lunar mission in this last decades.
STATE OF THE ART
In order to establish a first design concept, a benchmarking studio has been performed as a pre- study background work based on similar lunar missions which pursued similar objectives. In fact this state of the art assessment is a common practice before the concurrent engineering design fully starts, this way, early work on long-term subsystems can be gathered.
Among the variety of missions found, the following two have been found to be worth highlighted.
- Lunar Reconnaissance Orbiter, LRO: is a mission developed and managed as part of NASA’s Exploration Systems Mission Directorate. Particularly, LRO is a robotic mission aimed at creating a comprehensive atlas of the moon’s features and identifying available resources. This would be achieved in one year at low polar orbit around the moon, collecting detailed information about the lunar surface and environment.
Notice that although the extent of this mission is quite different from the one treated at the present design, some useful ideas for instance of possible further payload units are found. The differences are mainly due to the fact that this mission counts with a much greater scope, clearly exemplified by the dimensions of its space segment which besides the LRO itself, it also accounts for another spacecraft, LCROSS, which will directly determine if water ice occurs in an area of permanent shadow near the lunar poles. Therefore, orders of magnitudes of the main budgets treated along the project are not comparable
- SMART-1: ESA minisatellite technology mission to the Moon. It is the “first-ever" low-budget small mission for science at ESA. This mission goals were the study of Moon composition and the qualification of solar electric propulsion as a valid propulsion system. It was launched with Ariane onto a GTO, to first orbit the Earth in ever increasing ellipses, thrusting around perigee only to raise the apogee radius spiraling out from Earth/in toward Moon. Therefore the total transfer lasted for 13 months,
This mission orders of magnitude are much more comparable to the ones gathered at the present document. In fact, several ideas have been found in its solution design, such as Van Allen Belt escape, AOCS, optimisation of final mass to destination, Moon capture face, etc.
Notice that much other missions were consulted, for instance Change’4 Lunar Probe, which landed on the far-side of the Moon and carried out in-situ and rovering exploration. Nevertheless, for the sake of brevity, are not gathered at this section since in case of being used as references, they are properly mentioned along the document and referenced at section 6 REFERENCES.
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SYSTEM SEGMENTS AND MISSION GOALS
As aforementioned, the main purpose of the space system to be designed is to perform a fly-by pass through the Moon at an altitude lower than 100km with certain level of pointing accuracy ensured in order to capture the desired images accounting for a total of 10 MB of data,. Therefore, the three main segments conforming the system shall be design in accordance.
Firstly, in regard of the Ground Segment, this preliminary design only focusses on the selection of a station suitable with telecommunications S/S operational needs and which ensures mission requirements compliance, that is to say, that is limited to ESA Network stations.
Secondly, in relation to the Launcher Segment it is important to highlight that its analysis is out of the scope of this document. Therefore, just one, the Soyuz, is proposed based on its known capabilities to inject the designed nanosatellite into GTO at a free inclination. FIGURE 2 - Space System Segments Lastly, the Space Segment is the one deeply treated on the present document. It has been divided on two main aspects, the payload and the different subsystems. In relation to the former one, just an optical camera requested by the client has been included although it is possible to extent it with additional devices which may provide more accurate measurements of water existence, for instance a visible spectrometer to measure H2O vapour dissociation. This sort of equipment would require power, affect to the mass budget and imply rearrangement of internal S/C disposition. Nevertheless, it is expected that since the design has been done based on worst case scenarios and moreover wide margins have been included, in the case some additional equipment is desired to be placed, it would not detrimentally incur on subsystems design.
In regard to the subsystem composing the satellite, they are further explained at the following sections. Therefore, it is just necessary to highlight that they have been designed based on trade-off analysis ensuring that the following operating modes, corresponding to the different mission phases, are correctly tailored so that the expected performance is achieved.
- Commissioning phase. - Transfer phase: which leads to sun acquisition and pointing mode, transfer correction and thrust pointing mode. - Science/observation phase: corresponding to the fly-by operation itself. - Downlink data phase: once the fly-by is performed, Earth pointing is to be ensured to send science data back to Earth. This stage is highly related to the disposal phase of the mission.
Before entering into the requirements imposed to the system, it is relevant to briefly mention the phases in which the project has been divided in order to successfully overcome the design of the gathered nanosatellite.
1. Customer needs identification. 2. Requirement flow-down definition. 3. Mission concept exploration and trade-offs analysis between the most relevant design concept alternatives. 4. Preliminary design of different subsystems so that just one of the concepts is finally selected and properly sized accordingly to restrictions imposed by other subsystems. 5. Closure of preliminary design and drafting of the outcome PDR document,
Notice that phases 2, 3, and 4 are mainly characterised by its iterative nature. Therefore, it is of almost imperative need to develop engineering tools able to perform parametric studies and automate to the maximum possible extent the flexible response to changes, common in any CE processes.
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REQUIREMENTS FLOWDOWN
Requirements flow-down to engineering specifications is an essential step in modern systems design and development. It enhances the traceability management along the lifespan of the project and aids the identification and mitigation of design deviations. Therefore is a best practice that helps engineers derive or decompose functional and physical requirements from system level to the ones affecting their competencies.
In particular, the ones imposed by the client needs at the RFP for the Small-Satellite Mission to the Moon, are attached hereafter at Table 1.
TABLE 1 - Mission Requirements (High Level)
Notice that the Payload requirements are not included at system level together with the Mission requirements. On the contrary, they have been split into the Subsystems each of them involves.
Therefore, they are included at S/S level in the attached flow-down, as it can be checked at Appendix B – Requirements, where the final update of the flow-down list of derived subsystem requirements defined to fulfill the just exposed MR-xxx ones, is provided with further details for each of the considered Subsystems.
It is relevant to clarify the notation used along the project for the identification of the S/S requirements:
. Communication and Ground Segment S/S Requirements: CR-xxx. . Electric Power S/S Requirements: ER-xxx . Space Propulsion S/S Requirements: PR-xxx. . Attitude Determination and Control S/S Requirements: AR-xxx. . Mechanical and Structure S/S Requirements: MeR-xxx. . Thermal Control S/S Requirements: TR-xxx.
Where xxx stands for the numbering of each particular requirement within each of the S/S, starting from 001.
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3. SUBSYSTEMS ANALYSIS AND DESIGN
The core of the performed preliminary design is gathered at this section. In order to be able to properly understand its structure is important to notice three major facts:
- Each of the subsections begins with some clarifying statements as assumptions taken for the sizing or a brief summary of the main constrains affecting the treated S/S. In most of the cases those constraints are gathered at the low level requirements tables. - Particular trade-off criteria employed to discard the not suitable S/S design concepts are provided at Appendix D – Trade-Off Studies. For the majority of the subsystems, not only the weighted parameters but even the scoring resulting for the two main design options are also gathered at the Appendix. - Datasheets from which the relevant technical information about the treated components has been extracted are mentioned at the References Section. Some of the major component conforming the proposed design are moreover included at the Appendices section.
3.1 MISSION ANALYSIS
The mission analysis is based on the evaluation of the spacecraft trajectory on each of the mission phases in order to obtain an estimation of the Delta-V budget.
The concerning mission is composed of the following phases:
1) Transfer orbit from the GTO orbit in which the spacecraft is injected by the launcher (at a free inclination) to a distance equivalent to the distance between the Moon and the Earth. 2) B-plane targeting manoeuvers at the entrance of the Moon’s Sphere of Influence (SOI) 3) Fly-by over the Moon’s surface at an altitude lower or equal to 100km 4) Exit from the Moon’s SOI and disposal phase.
DELTA-V BUDGET For the calculation of the Delta-V however, the most significant phase is by far is the transfer leg, given that the Transfer Corrections maneuvers required during the transfer and B-plane targeting and the disposal trajectory are minor with respect to the former, and the fly-by is only driven by the Moon’s gravity field.
Therefore, in order to obtain an estimation of the Delta-V budget, the following assumptions have been considered:
- Only the transfer maneuver from the GTO to a distance equivalent to the Moon orbital radius will be assesses for the estimation.
- No perturbations will be considered for simplicity in this very preliminary study. It is expected that the drag contribution would provoke a non-conservative effect on the calculations during the first orbits of the trajectory, but the third body perturbation of the Moon could help to counteract it to some extent.
- A 10% Delta-V margin will be considered. This margin accounts for drag perturbation at the perigee of the transfer orbit, for the B-plane targeting maneuvers for injection in the Lunar SOI and other Trajectory Correction Maneuvers (TCM) for contingencies. Notice that Sun and Moon will provoke perturbations in the orbit inclination, but they can be counteracted by choosing properly the injection in the GTO, and hence the expected Delta V is negligible. Moreover, as aforementioned, the flyby maneuver itself will be assumed to not require any propellant since it will be driven by the Moon’s gravitational field in a hyperbolic trajectory.
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Finally, after the flyby, the spacecraft will be assumed to be lost in space (positive flyby), and hence no additional disposal Delta V will be required.
- In principle, attitude control will not be a requirement of the main propulsion system. Nevertheless, the previous Delta-V margin is expected to be sufficiently large in case this possibility is considered in further analysis of the project.
- No maneuvers to change orbital plane have been considered since the launcher provides free inclination GTO.
After acknowledging these assumptions, it is worth mentioning that there is a strong co- dependence between the mission analysis and the propulsion system analysis as the latter will have a large impact on the transfer trajectory of the spacecraft. After all, the Delta-V budget and the mass budget are directly related by Tsiolkovsky equation.
Δ푉 푚푒푛푑 = 푚𝑖푛𝑖 · exp(− ) 퐼푠푝푔0
Please, refer to the Propulsion subsystem section for the rationale of the selection of the two representative thrusters for the chemical and electrical propulsion types: hydrazine monopropellant liquid rocket, and IFM-NANO electrospray thruster, respectively.
For the chemical propulsion alternative, the Delta V estimation is performed using a transfer similar to a Hohmann but without recircularization, that is, just one impulse transfer at the GTO perigee were the energy of the S/C is maximum and therefore the required Delta-V to the Moon orbital radius, minimum: μ(1 + e ) μ(1 + e ) √ t √ GTO Δ푉푐ℎ푒푚𝑖푐푎푙 = 푉퐻표ℎ푚푎푛푛푝푒푟푖푔푒푒 − 푉퐺푇푂푝푒푟푖푔푒푒 = − at(1 − et) aGTO(1 − eGTO)
Where 휇 is the Earth’s gravitational constant; and 푎푡, 푎퐺푇푂, 푒푡, 푒퐺푇푂 are the semi-major axis and eccentricity of the transfer and GTO orbits, respectively.
It is important to notice that the Delta-V calculation in this occasion is independent from the spacecraft mass, since it only depends on the orbital parameters.
For the electric propulsion alternative, the Delta-V estimation is slightly more complicated since the initial orbit of the spacecraft (GTO) is highly elliptic. Hence, the simple calculation of the difference in velocity between the initial and final circular orbits is not enough accurate. Therefore, for this scenario, the transfer orbit has been propagated using MATLAB Spice module (Cowell propagator), assuming a constant mass flow rate and a thrust contribution to the acceleration given by:
푇 푇 푚̇ 푆/퐶 = − ; 푎푐푐푡ℎ푟푢푠푡 = 퐼푠푝푔0 푚푆/퐶
This way, knowing the final mass of the satellite, it is possible to make an estimation of the propellant 푚푒푛푑 mass consumed, and then, apply Tsiolkovsky equation to obtain the Delta-V: Δ푉 = −퐼푠푝푔0 ln 푚푖푛푖
Nevertheless, this calculation imposes a very important restriction since it is necessary to input a value for the initial mass of the spacecraft in order to perform the orbit propagation.
Therefore, since at this initial state of the design, no data on the spacecraft total mass is available, calculations are to be performed by means of a parametric analysis assuming values for the spacecraft mass between 10 and 54 kg.
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Having said this, Figure 3 presents the Delta-V results for the chemical and electrical propulsion alternatives, accounting both for the 10% abovementioned margin.
From this figure, it must be firstly noticed that the calculation for the electric propulsion system is only available up to a spacecraft mass of 18kg. This value stands for the maximum mass for which the mission duration is compliant with the 5-years requirement (MR-008).
FIGURE 3 – Delta-V Budget
From the previous graph, it can also be observed how the Delta-V of the electric propulsion is approximately 2km/s larger than for the chemical thruster. In principle, this would imply a larger amount of fuel consumed since the ratio 푚푒푛푑/푚𝑖푛𝑖 is smaller. Nevertheless, the specific impulse of the electric thruster is one order of magnitude larger than the one of the hydrazine, and therefore the latter statement turns out to be false.
Moreover, the final-to-initial mass ratio is far more determinant than the Delta-V budget to reach a conclusion regarding the best propulsion alternative. Consequently, please refer to Space propulsion Subsystem section for the reasoning on why the electric thruster has been selected over the chemical thruster for the propulsive alternative for the concerning spacecraft.
Finally, after obtaining the final wet mass of the S/C, the Delta-V is found to be 2.74 km/s.
SPACECRAFT TRAJECTORY It is important to mention that the goal of this preliminary design is not to fully determine the spacecraft trajectory all along the mission but provide a preliminary trajectory which serves so as to size all the required subsystems of the spacecraft.
Having said this, the spacecraft trajectory will be treated separately for the transfer and fly-by maneuvers leaving the targeting analysis as an open subject for further determination during the detailed design phase of the project.
The main reason for this decision is the lack of specification of two aspects that normally are determinant to perform a launching window analysis: the launching phase of the project, which indeed is out of the scope of this preliminary analysis; and lack of specification of the Moon’s region where the fly-by should be performed.
In regard to the launching phase, the only three known facts about the launcher is that it will have to be launched from an ESA station and inject the spacecraft on a GTO orbit at a free inclination.
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However, the launching site has two main characteristics that affect the targeting analysis: the climate of the region (which normally sets the time of the year in which the ambient temperature is suitable to perform the launching), and its latitude (which affect the maximum inclination of the orbit that can be reached at a “free cost”).
On the other hand, since there is no clear specification on the mission requirements about the region of the moon in which the fly-by shall be performed, there are no constraints on the transfer and fly- by trajectories to this regard.
Hence, there is a whole spectrum of possibilities which would have to be investigated to select the optimal launching period of the year, inclination of the GTO orbit and region of the fly-by.
Once, the rationale for the assumptions for the trajectory analysis has been introduced, the specific trajectories for the transfer and fly-by maneuvers is to be presented. Please, bear in mind that these trajectories have been calculated for the final spacecraft mass resulting from the refined design process of all the others spacecraft subsystem: 8.57 kg.
Transfer trajectory
For the transfer orbit, the specific assumptions were the following:
- Arbitrary ephemerides (since no targeting analysis is considered).
- Transfer maneuver goes from the GTO to a distance equal to 384400 km (Moon orbital radius about the Earth)
- Orbit inclination: from the above stated, this value could be selected at will since there are no restrictions from the launching capabilities or the fly-by location in the moon. Therefore, it was select to be 0º with respect to the ecliptic plane, given that this is considered to be the most restrictive condition for the dimensioning of the batteries of the spacecraft power subsystem. However, notice that selecting a larger inclination would help to minimise the occurrence and the duration of eclipses, therefore minimizing the size of the power batteries, and the time spent under the radiation of the Van Allen belts (important for sizing the spacecraft radiation protection). Moreover, a higher inclination of the orbit would also help the fly-by to take place nearer the Moon’s poles, where the probability of encountering water is larger. This refined trajectory study is left open for the detailed definition project phase.
- Portion of the orbit (measured in true anomaly angle) in which the thruster is providing an acceleration to the spacecraft (Requirement PR-004 in Appendix B – Requirements): angle of +/-160º measured from the perigee of the orbit. In terms of time, this would correspond approximately to a 65% of the orbital period. Theoretically, the choice of performing the firing around the perigee of the orbit allows to concentrate the thrust force on increasing the apogee altitude, allowing to reach faster the SOI of the Moon. Nevertheless, it would have been interesting to perform an analysis where the thruster firings were performed close to the apogee to see the effect on the fly-by and disposal trajectories.
- No orbital perturbations have been considered for the transfer orbit. Drag and solar radiation pressure can be considered small, given that the first only acts along the perigee of the orbit, and the spacecraft cross-section is small compared to its mass for the solar pressure to provoke significant perturbation forces. Nevertheless, the third body perturbation could signify an important factor since it would act all along the orbit (not only in the perigee). However, this contribution will mostly help to reduce the transfer time since the spacecraft will be driven by the Moon’s gravity once it enters its SOI. Therefore, by not considering this perturbation, the computation of the transfer duration and the number of revolutions (which are most important for the calculation of the accumulated momentum in the dimensioning of the reaction wheels of the AOCS subsystem and the number of eclipses to be taken into
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account for the batteries dimensioning in the power subsystem) will provide more conservative values.
The results of the transfer trajectory are available in Figure 4.
FIGURE 4 - Transfer Trajectory
Even though no targeting analysis is being performed, it seems appropriate to assume that the launch window would be such that the position of the Moon favors the capture of the spacecraft by the Moon SOI (i.e. apogee of the elliptical spiral trajectory for the case of electric propulsion), as represented in the previous picture.
Figure 4 corroborates that the transfer duration calculation is conservative since the last 3 orbits lie inside the Moon sphere of influence, implying that the B-plane targeting could have been performed in a previous intermediate orbit. Moreover, the value of the 65% for the thruster firing time is also conservative. However, since the transfer duration (approximately 17 months) is far from the maximum mission duration of 5 years there is no concerns to this regard.
Another aspect that can be extracted from the transfer orbit is the time spend under the influence of the radiation of the Van Allen belts. This is simply calculated by measuring the time periods in which the spacecraft altitude lies between 500 and 5000 km (inner Van Allen belt) or between 15000 and 58000 km (outer Van Allen belt). The period spent under this radiation is about 7.33 months, which shall be important when determining the thickness of the spacecraft casing in the mechanical subsystem design.
Fly-by trajectory
Analogously as for the transfer orbit, this trajectory will be computed using MATLAB Spice module for the hyperbolic orbit propagation, this time in the Moon centered reference frame. As before, no perturbations will be considered although in this case their effects are even more negligible since the duration of the fly-by (order of hours) is very small in comparison with the transfer orbit (months).
FIGURE 5 - Fly-By Trajectory
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Moreover, recall that, since there is no explicit requirement stating the region of the Moon’s surface to be studied during the flyby, it has been considered to occur at the Moon’s equator,
The duration of the fly-by is obtained to be 29.9 hours, measured from entrance to exit of the Moon’s SOI. Nevertheless, the period in which the payload camera is activated is of the order to minutes. Moreover, the angle formed by the entry and exit velocity vectors is 151º, meaning that the angle formed by the two arms of the hyperbolic trajectory is 29º.
Disposal trajectory
The disposal trajectory is highly dependent on the fly-by trajectory, as it determines whether the fly- by is positive, negative or neutral.
In this case, the fly-by trajectory is so narrow (29º) that the fly-by will almost always be positive,
gaining velocity with respect to the Earth, as represented in Figure 6, where 푉∞1, 푉∞2 represent the entrance and exit of the Lunar SOI velocity vectors with respect to the Moon, and 푉1, 푉2 with respect to the ECI reference frame. The angle 훼 is the SOI injection angle, which can be partially tailored during the target analysis to be performed in the Detail Design Phase of the project.
FIGURE 6- Fly-By Velocity Vectors
FIGURE 7 – Disposal Trajectory
Figure 7 shows the different alternatives for the disposal trajectory for three different values of the SOI injection angle 훼. Notice that these trajectories have been propagated for a period of 7 days, as this is the required period estimated by the Communications Engineer to transmit the payload data after the fly-by trajectory. As it can be seen, both red (57 º) and magenta (28 º) trajectories are suitable, as they pass close to the Earth during the escape trajectory.
After analysing all the mission phases, the total mission duration is found out to be 18.4 months, which fulfils the mission requirement MR-008 with a comfortable margin.
As a final comment, the targeting analysis to be performed in the detailed design phase shall take into account that the camera does not point towards the sun while performing the fly-by, and that the spacecraft does not project a shadow over the Moon’s surface being filmed. Nevertheless, these are easy requirements to fulfill by selecting a proper launch window and fly-by trajectory inclination.
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3.2 SYSTEMS OPERATIONS MODES
Commissioning Mode: The first 15 days orbiting around the Earth are used to ensure that all the subsystems work correctly, for instance, try the safe mode where the solar arrays point towards the sun to charge the batteries.
Sun acquisition Mode: Initially at CubeSat deployment for power positive control; It would use RW and sensors, with gyroless submode for fail safe operation. [1]
Safe mode: Is the recovery mode for emergency situations. Its purpose is to ensure power generation and ground communications using a minimum of on board resources.
Sun pointing Mode: The satellite spends almost all the mission time travelling from Earth to Moon. The requirement associated to this phase is to ensure pointing towards the sun in order to guarantee correct solar arrays performance. On the other hand, according to requirement AR-001, the payload must never point directly toward the Sun, therefore camera shall be hidden during this mode.
For this purpose, the attitude control will not play any role as the solar panels can be oriented using mechanical actuators to point towards the Sun. Under safe mode operation, six sun sensors are used for spacecraft and solar array pointing to ensure power generation [4]. By measuring the current produced by 3 sensors that share an apex, the sun vector can be calculated using the cosine law.
Mission Transfer Correction Mode: External torques will destabilise the satellite. Correction mode shall ensure these torques are compensated during the entire mission duration.
Thruster Pointing Mode: In order to make effective the thrust provided by the selected propulsive subsystem, the thrust line has to point tangentially to the orbital desired trajectory. As the satellite orbits, the thrust line is lost so the AOCS is required to act proving the correct attitude which allows to thrust in the required direction.
Earth Pointing: To allow send data back to the Earth after the flyby so that the antenna (X axis body frame) points towards the G/S.
Science Observation Mode: This mode allows the satellite to point towards the centre of the Moon during the fly-by in order to guarantee the required pointing accuracy. Three-axis stabilization and correct sensor selection shall allow to meet the accuracy levels. The main ones are: RW, star trackers and IMU, to provide inertial, sun, and nadir attitude control as well as slew manoeuvres. [1]
3.3 SPACE PROPULSION SUBSYSTEM
In regard to the propulsion subsystem, there are three key mission requirements which may help to automatically discard some types of space propulsion: the payload mass, the mission duration and the type of trajectory maneuvers. Given the low value of the payload mass (1 kg), the relatively wide maximum mission duration (5 years) and the lack of need for impulsive maneuvers (such as changes in orbital plane), both chemical and electric propulsion seem to be suitable for the mission, and therefore a deeper trade-off analysis must be performed.
It is worth mentioning that, other types of propulsion as nuclear propulsion or propellantless (e.g. solar sails) have not been taken into account mainly by having a lower maturity state, but also due to social concerns (nuclear) or extremely low thrust force (solar sails).
CHEMICAL VS. ELECTRICAL PROPULSION On one hand, regarding chemical propulsion, there are three main possibilities: solid, liquid or hybrid propellant. Nevertheless, combustion in solid rockets cannot be stopped once ignited, which may be desirable for some applications, such as rocketry, but it does not fulfill the requirements of the
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Lunar flyby; and hybrid propellants are still under investigation, since it is a relatively new concept and they impose some new constraints with respect to the other two. Therefore, liquid chemical propulsion shall be considered for the trade-off analysis.
Moreover, liquid propellant rockets may be monopropellant or bipropellant. However, given the limited spacecraft volume (27U maximum) and the higher complexity of the bi-propellant systems (separate tanks for each reactant, distribution and mixing process, etc.), a monopropellant rocket will be evaluated.
The specific thruster was selected to be a 1N-thrust (monopropellant) hydrazine (N2H4) thruster due to is extended use and its relatively high specific impulse (~ 220 s).[25]
On the other hand, for the selection of the representative electric propulsion engine, there is a large variety of propulsive concepts although they can be integrated in three main groups: electrothermal (resistojets, arcjets), electrostatic (FEEP, ion thrusters) and electromagnetic pulsed (PPT) and non- pulse (HET) thrusters.
From all these possibilities, some can be directly discarded due to their low performance or demanding requirements. First, electrothermal thrusters’ performance (thrust, Isp) is limited by the melting temperature of the heating element used. HET thrusters normally require an extremely high power input. Pulse Plasma Thrusters (PPT), in spite of requiring small inputs of power (~10W), have issues regarding both the non-uniform ablation of the solid fuel which result in low efficiency, and the high electrode erosion. For these reasons, it seems that Field Emission Electric Propulsion (FEEP) (also referred as to electrospray propulsion) and Radio-Frequency Ion Thrusters (RIT) are the two most promising alternatives for electric propulsion for this Lunar mission.
Nonetheless, FEEP thrusters require less power than the RIT thrusters (at the expense of a slightly lower thrust and Isp). Hence, since no preliminary design for the electric and power subsystem was available at this early stage of the design, and in order to minimize the power demands of the propulsion subsystem, it was decided to select the FEEP thruster as the alternative to perform the trade-off with the liquid chemical thruster.
The particular FEEP thruster selected was the IFM-NANO thruster [23] whose operational envelope allows feeding power values from 20 to 40W for thrust levels between 0.2 and 0.4mN.
Once, the two alternatives for the trade-off analysis are selected, it is possible to obtain a first guess for the propellant mass budget.
As aforementioned in the mission analysis section, the Delta-V budget is closely linked to the propellant mass budget by Tsiolkovsky equation. In fact, recalling the assumptions for the Delta V calculation presented in said section, for the chemical propulsion system the propellant mass will be obtained from the Delta V calculation (2.49 km/s) and using Tsiolkovsky equation: whereas for the electric propulsion case, the Delta V was inferred from the propellant mass budget resulting from the orbit propagation using MATLAB Spice.
Therefore, for chemical propulsion the propellant mass will be obtained from the value of Delta-V provided in the mission analysis section (0.69 km/s) and using Tsiolkovsky equation, whereas the propellant mass for the electric propulsion alternative will be obtained from the final mass of the spacecraft provided by the Keplerian orbit propagation: 푚푓푢푒푙 = 푚𝑖푛𝑖 − 푚푒푛푑(퐶표푤푒푙푙).
Nonetheless, another extremely important factor for this trade-off analysis is the mass of the propulsion subsystem itself.
The assumptions for the chemical propulsion subsystem are:
- Pressurization system: Blow-down pressurization system (to avoid inclusion of a second tank for the pressurant gas) using Helium as the pressurant.
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- Fuel storage: Titanium tank sized to withstand pressures between 5.5 and 22 bars (inlet pressure range of the thruster) and accounting for a 200% margin (burst pressure) for the walls thickness sizing.
- Hydrazine thruster mass: 0.3kg [25]. Number of thruster calculated from the single burn lifetime of the thruster and the nominal fuel mass flow rate.
Regarding now the electric propulsion subsystem, its assumptions in relation to the thruster are:
- Operational point of the thruster inside the envelope: Thrust=0.35mN, Isp=3000; total impulse=7350Ns; nominal power=32W. [24]
- Number of thrusters needed: computed from the transfer trajectory duration and the total 푑푢푟푎푡𝑖표푛 impulse given in the envelope as follows: 푛 = 푡ℎ푟푢푠푡푒푟 푇표푡푎푙𝑖푚푝푢푙푠푒/푇ℎ푟푢푠푡
For the assumptions of the orbit transfer considered, please refer to the mission analysis section of above.
Hence, Figure 8 presents the results of the propellant mass budget for the trade-off analysis between chemical and electric propulsion. The two showed dotted lines represent the available empty mass for the rest of the spacecraft subsystems. Notice that the jumps in the electric P.S. scenario are due to the need of introducing a new thruster due to lifetime issues.
FIGURE 8 - Propellant Mass Budget
From these results, it seems that electric propulsion is only suitable for small spacecraft mass, whereas chemical propulsion would only be efficient for the largest range of masses considered up to the limit imposed by the requirement MR-007.
In principle, it seems coherent to select the electric propulsion alternative since for the same values of the S/C wet mass, it allows a higher mass for the remaining subsystems. This last fact will imply a significant cost reduction for the project budget, especially regarding launching-related costs, which even though is out of the scope of this preliminary design, it is an important selection criteria to be considered in real projects.
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However, some other secondary disadvantages of electric propulsion shall also be considered:
- While chemical propulsion is said to be energy limited (mostly due to the propellant properties), electric propulsion is power limited. For electric thrusters, the power budget is higher, although this might be alleviated by firing the thrusters when the rest of subsystems are off or in stand-by conditions. The power supply during eclipses is critical given that the sizing of the battery would have to be impractically large. This can be rather restrictive for LEO orbits, however, the Lunar transfer trajectory can be selected in further stages of the project to avoid eclipses as much as possible by selecting adequate values of the argument of perigee and the inclination for the orbit.
- Also, plasma plume may interact with other spacecraft components provoking: S/C charging, sputtering or contamination of other surfaces (solar panels), or interaction of the plasma during the spacecraft communications.
In light of the above, it seems appropriate to select the electric thruster as the propulsive alternative for the mission given its higher efficiency and cost reduction in comparison to chemical thrusters.
REFINEMENT PROCESS OF THE ELECTRICAL THRUSTER As it has been stated previously, the decision to select an electric thruster imposes two major restrictions on the rest of spacecraft subsystems regarding mass and power. Moreover, there is an important propulsion requirement which has not been implemented up to now regarding the inoperative time of the propulsion system.
Requirement PR-004 imposed the thruster to be switched OFF at least 30% of the mission duration, which in rough numbers would imply a 30% increase in the total mission duration (being all other parameters fixed). Therefore, this requirement constraints even more the maximum total mass of the spacecraft since in order to fulfill the 5 years of maximum mission duration, the mass would have to be reduced assuming the thrust force remains the same. Notice however, that the implications of this requirement regarding the thruster lifetime are minor since the total time in which the thruster is operative would be mainly the same.
Therefore, the previous study for the IFM-NANO was performed again accounting for this inoperative time. Moreover, several points of the operational envelope were evaluated in order to study the effect of varying the thrust, Isp, and total impulse. Additionally, other alternative FEEP thrusters with different performances were also considered (TILE5000 [24] and BET-1mn [22]). The study was performed also for values of the spacecraft wet mass ranging from 7 to 15 kg, nonetheless, only the results corresponding to 9 kg of wet mass are provided in Table 2 for the sake of shortness:
Total Transfer Dry Thrust Thruster Power Delta-V Number Of Thruster Isp (S) Impulse Duration Mass (mN) Mass (kg) (W) (km/S) Thrusters (Ns) (Months) 1(Kg) IFM-NANO (56) 0.27 4000 9800 0.67 30 2.74 (4) 48.5 (1) 4 (2) 5.71 (5) IFM-NANO (54) 0.3 3000 7350 0.67 28 2.74 (4) 43.2 (2) 5 (2) 4.85 (5) IFM-NANO (50) 0.35 3000 7350 0.67 32 2.74 (4) 37.1 (2) 5 (2) 4.85 (5) BET-1MN (0) 0.7 800 605 1.15 15 2.74 (4) 16.5 (4) 50 (0) - (0) TILE-5000 (0) 1.5 1500 4900 1.4 25 2.79 (5) 8.36 (5) 7 (1) - (0)
TABLE 2- FEEP Thrusters Trade-Off for Wet Mass of 9kg
1 From now on this dry mass will be considered not to include the mass of the propulsion subsystem, to be addressed separately.
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From the previous results it is clear that the two thrusters providing a larger amount of thrust, TILE5000 and the BET-1mN, have an insufficient total impulse (i.e. lifetime) which results in an extremely large number of thrusters that indeed leaves few available mass to accommodate the rest of the subsystems. On the other hand, the IFM-NANO shows a good performance with a transfer duration inside the mission requirements and a comfortable available mass for the rest of the system, especially for the first considered operational point of the envelope (0.27 mN thrust).
Nonetheless, at this stage of the preliminary design, it was expected that the total volume of the rest of the subsystems would range between 2 and 3 U, whereas the propulsion system alone would require approximately 5U. Since this implies a challenge for the spacecraft mechanical design regarding the distribution of all the subsystems, it was decided to further research the field of gird ion thrusters looking to reduce the Propulsion subsystem volume.
As aforementioned, ion thrusters could also be viable for the mission since they offer slightly higher values of thrust although at a higher power cost with respect to FEEP. This larger power budget imposes a constraint since at this state of the design the power subsystem is being designed to offer power values up to 30W for the propulsion subsystem.
Nevertheless, an unprecedented model of ion thruster, Busek BIT-3 [21],, has been found to exhibit an extremely good performance. As it can be seen at Table 3, this new thruster provides 1.4mN thrust with an Isp of 3000 s. Moreover, the lifetime is only limited by the ion grids lifetime (20000h = 27.8 months), so in principle, one thruster could be enough to perform the whole mission.
Additionally, this thruster features three characteristics which are major advantages for other spacecraft subsystems as well [26]:
- The thruster is prepared to use Iodine instead of Xenon, allowing storage in solid state (i.e. higher density) and at a much lower pressure than Xenon, which is also more costly.
- The thruster include a 2-axes, +/-10º gimbal which allows to direct the thruster, in order to create a momentum, and avoiding the need to include a RCS system for desaturating the reaction wheels of the AOCS subsystem. Please refer to the section of this subsystem for further details.
- The thruster has already being certified for space flight, and it is currently being used by the LunarCube spacecraft launched in 2018. [1]
FIGURE 9 - BUSEK BIT 3 Thruster.
Nevertheless, the major disadvantage is that, even though the operational power range covers from 20W to 70W, its nominal power is 60W, meaning that a lower power supply would imply a large
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decrease on fuel efficiency [19]. This can be corroborated by Table 3 in which the results of two operational points of the BIT-3 are provided assuming a wet mass of 9 kg.
Ion Fuel Transfer Number Dry Thrust Isp grids Thruster Power Δ-V Thruster utiliza- duration of mass (mN) (s) lifetime mass (kg) (W) (km/s) tion(%) (months) thrusters (kg) (h) BIT-3 (0) 1.4 3000 20000 1.5 60 (0) 70 2.83 (3) 9.5 (5) 1 (5) 6.32 (5) BIT-3 (62) 0.7 2000 20000 1.5 35 (1) 40 2.76 (3) 18.2 (4) 1 (5) 4.54 (4)
TABLE 3 - Bit-3 Results for 9kg Wet Mass
As it can be observed, by reducing the power, the fuel utilization efficiency drops by a 30%, causing the need of carrying more fuel to obtain the same thrust. Comparing the results of BIT-3 in Table 3 for 35W with the results for the IFM-NANO in Table 2, it can be seen that the available empty mass for the rest of the subsystems is lower in the former.
Nevertheless, even though the mass of the propulsion system is larger for the BIT-3, its volume has been reduced from 5U to 2U, one for the thruster, cathode, gimbal and PPU, and another for the fuel tank. This great reduction in volume, together with the advantages introduced by the gimbal, the easy storage of the fuel and the certification status of the thruster, have been the key factors to select the Busek BIT-3 as the thruster for the LASAR CubeSat.
It is important to notice, that in order to increase the power budget It is important to notice, that in order to increase the power Budget of the thruster from 30W, required by the previous IFM-NANO thruster, to 40W for the BIT-3 (35W for the thruster and an additional 5W for a resistor to vaporize the Iodine from its solid state) [19], it is required to adopt the assumption that the thruster will not be operating during the trajectory eclipses so as to avoid over-dimensioning the power batteries. According to the mission analysis section, this assumption is very reasonable mainly due to two factors: the inclination of the orbit can be selected to minimize the occurrence and duration of the eclipses; and the thruster is inoperative during 65% of the orbit, more than enough to cover the eclipse duration.
Finally, once the dimensioning of the rest of the spacecraft subsystems has been refined, with a global mass of 8.57 kg, the final propulsion subsystem mass budget can be determined as portrayed at Table 4:
Transfer Propulsion Total Delta-V Propellant Dry Mass Thruster Duration Subsystem Mass Power (W) (km/s) Mass (kg) (kg) (Months) (kg) BIT-3 40 2.74 17.2 2.79 1.5 4.28
TABLE 4 - Propellant Mass Budget and Propulsion data
As a final comment, given that the thruster will be used for the desaturation of the wheels, it is important to verify that the fuel consumed to this end (15g)2 is already covered by the 10% Delta-V margin. Indeed this margin corresponds to approximately 95g fuel, and recalling from the mission analysis that neither the fly-by nor the disposal will required any Delta V, the remaining 80g are expected to be sufficient to counteract the drag perturbations of the transfer orbit during the first passes by the perigee and perform the required TCM for the Lunar SOI injection.
The detailed calculation of the propellant required for the TCM is supposed to be performed in the Detail Design Phase of the project once the targeting analysis and the complete determination of the transfer orbit have been performed.
2 Please, refer to the AOCS section for the details on this calculation.
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3.4 ATTITUDE, DETERMINATION AND CONTROL SUBSYSTEM
This section presents the trade-off and selection of the ADCS components as well as their sizing to control the nanosatellite attitude for each of the presented mission operations mode, throughout the entire mission duration.
OBJECTIVES OF THE ADCS - Find the current attitude of the SC: Which provides S/C attitude knowledge to support mission objectives. - Reorient S/C as required - Stabilise S/C attitude: Providing rate stabilization and pointing for payload, power, communication and thermal subsystems during normal and safe operations.
For most of the mission, a good initial approximation for the satellite attitude is the Torque-Free Satellite motion. In absence of external torques, the total angular momentum ℎ is conserved. The Euler equations of motion for torque free satellite motion are provided hereafter, being the equilibrium guaranteed if these are fulfilled.
퐼푥휔푥̇ − (퐼푦 − 퐼푧)휔푦휔푧 = 0퐼푦휔푦̇ − (퐼푧 − 퐼푥)휔푥휔푧 = 0퐼푧휔푧̇ − (퐼푥 − 퐼푦)휔푥휔푦 = 0
However, external torques are applied during the mission, so actuators need to be used to compensate them as well as to ensure the correct attitude that allows to meet the mission requirements.
Therefore, the main elements playing a role within this subsystem are:
- Sensors: To measure the attitude. - Actuators: To exert torques on the satellite structure in order to change its attitude. - Computer and algorithms: To process the collected data and decide the course of action. - Ground Control: To receive the measured attitude and input commands to the ADCS.
DISTURBANCE TORQUE ENVIRONMENT The most important sources of torque that would lead to a loss of attitude control during the mission are to be identified and analysed in this section, enabling the AOCS subsystem selection and sizing.
Gravity Gradient:
The gravity gradient consists on the variation rate of gravity force along the vertical. This effect can destabilise the S/C and has a great importance for large structures since it is highly dependent on the inertia of the spacecraft as well as on the orbit altitude. This gravity torque can be modelled as:
푀푔 = ∫ 휌^푔푑푚
For what the resulting equations of motion, where 푟 = (푋, 푌, 푍) stands for the position vector of the S/C centre of mass measured from Earth centre, are:
3퐺푀 3퐺푀 퐼 휔̇ − (퐼 − 퐼 )휔 휔 = 푌푍(퐼 − 퐼 ) 퐼 휔̇ − (퐼 − 퐼 )휔 휔 = 푍푋(퐼 − 퐼 ) 푥 푥 푦 푧 푦 푧 푟5 푧 푦 푦 푦 푧 푥 푥 푧 푟5 푥 푧
3퐺푀 퐼 휔̇ − (퐼 − 퐼 )휔 휔 = 푋푌(퐼 − 퐼 ) 푧 푧 푥 푦 푥 푦 푟5 푦 푥
The worst case scenario for the gravity gradient is therefore observed when the difference of moments of inertia is maximum. This, depends on the Spacecraft geometry so as a first approximation, it has been considered that:
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3퐺푀 푇 = (퐼 − 퐼 ) sin(2휃) 퐺푟푎푣𝑖푡푦 2푟3 푦 푥
Where 휃 is the maximum deviation of the Z-axis from local vertical in radians. Notice that the most conservative value for 휃 has been chosen.
Solar Radiation Pressure:
Solar Radiation Pressure strongly depends on the attitude strategy as well as on the properties of the S/C surface which is illuminated. A surface can either be transparent, absorbent or reflective, although most surfaces are a combination of the three. In general, solar arrays are treated as absorbers and the spacecraft body as a reflector.
In particular, this external perturbation is mainly driven by the S/C geometry, the surface reflectivity and the cg location. But since they are not finally defined a worst-case solar radiation torque is
퐹 푇 = 퐹(푐 − 푐 ) 퐹 = 푠 퐴 (1 + 푞)cos(퐼) 푠푝 푝푠 푔 푐 푠
2 8 Being 퐹푠 the solar constant,1.37푤푚 ; c the speed of light, 310 푚/푠 ; 퐴푠 the surface area; cps the location of the centre of solar pressure; 푐푔 the centre of gravity; 푞 the reflectance factor (ranging from 0 to 1, 0.6 will be used), and 퐼 the angle of incidence of the Sun. Again the worst case scenario has been selected by assuming퐼 = 0푑푒푔.
Aerodynamic Drag:
The torque due to atmospheric drag is to be evaluated as:
1 푇 = 퐹(푐 − 푐 ) 퐹 = 휌푉2퐴 ∗ 퐶 푎 푝푎 푔 2 푠 푑
Being 퐶푑 the drag coefficient, which typical values are 2-2.5, 휌 the atmospheric density, 퐴푠 the surface area, 푉 the S/C velocity, 푐푝푎 the centre of aerodynamic pressure and 푐푔 the center of gravity.
The worst case scenario in this occasion is the one at lower altitudes, leading to higher density and high velocity. Therefore, this is highly related to the passes close to the Earth. A drag coefficient equal to 2.5 will be considered.
Van Allen Radiation Belts:
The issues caused by the Van Allen Belts lie on the high amount of charged particles they enclose. In case of Apollo missions, the solution was to minimize the time of exposure to those particles as well as to avoid regions with higher charged particles density. Nevertheless, in the present mission, as electric propulsion is being used, long time exposure cannot be avoided. [3]
However, it is not the first time an electric propulsion spacecraft travels through the Van Allen Belts. SMART-1 Mission was operated in electric propulsion mode almost continuously. Regarding the lessons learnt about AOCS subsystem, during SMART-1 Mission a star tracker was lost, but the back- up system allowed to continue obtaining good attitude determination. These problems were related to radiation damages, soft protons, severe solar storms, but mainly to longer periods of time cruising through the inner Van Allen belts. The rest of subsystem devices (gyros, reaction wheels, gimbal and thrusters) did not suffered any problem due to the high radiation levels.
Keeping in mind this lesson learnt, it will be defined a back-up system for sensors since they are the AOCS elements suffering more due to the Van Allen Belts passes. It has been considered that the extra perturbance radiation pressure torques are covered by the high selected margins in the present PDR stage.
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Magnetic Torque:
Magnetic torque is caused by the interaction between the geomagnetic field 퐵 and the satellite residual magnetic fields. It can be modelled as:
푇푚 = 퐷 ∙ 퐵
Where 퐷 is the residual dipole of the vehicle in 푎푚푝 ∙ 푡푢푚 ∙ 푚2, and 퐵 can be approximated as 2푀/푟3 for a polar orbit to half that at the equator. 푀 is the magnetic moment of the Earth, equal to 7.96 ∙ 1015푡푒푠푙푎푚3, and 푟 is the radius from the Earth centre to the S/C in meters.
As it can be observed, this perturbation is highly dependent on the orbit altitude, inclination, and the residual spacecraft magnetic dipole.
Other environmental or operational torques
- Uncertainty in Centre of Gravity (cg): It cannot be fully defined, which could lead to unbalanced torques during firing of the thruster. - Rotating Machinery: Depends on the S/C design which is still on preliminary phase. - Liquid Sloshing: Torques due to fluid motion and variation in centre of mass. Not being such relevant as solid propellant is being used. - Dynamics of flexible bodies: Oscillatory resonance at bending frequencies. - Thermal Shocks on Flexible Appendages: Attitude disturbances when entering leaving eclipses.
Since these specifications are much more difficult to estimate at a preliminary design phase, high margins have been applied to ensure proper operation.
DESIGN SOLUTION FOR THE ACDS The sizing of the AOCS system shall ensure the satellite control in all the operations modes. As in other subsystems, the design is to be performed according to the most restrictive conditions expected during the mission.
It is important to bear in mind that as electric propulsion is being used, the mission time is increased and the perturbance torques, which are small in magnitude, are accumulated during a big amount of time. Thus, making desaturation to play an important role in case of selecting reaction wheels as ACDS system. The need of desaturation implies to design a RCS system directly implying an increase in mass of the S/C. At least three wheels are required with their spin axes placed in a non-coplanar manner. Often, a fourth redundant wheel is carried in case one of the three primaries fails. [2] On the other hand, in case of using any other type of actuators, the number of pulses will be high, and influencing therefore their lifetime, which needs to be taken into consideration.
Regarding the way to measure the accumulated angular momentum and the maximum torque produced by perturbations, once the evolution of the radial position with respect to Earth at every instant is known from the trajectory provided at section 3.3 the control and perturbance torques can be completely defined.
Gravity Gradient & Magnetic Torque: Knowing the evolution of the radial position with time, it is possible to obtain an estimation of the required control torque to correct this perturbation and thus achieve the required attitude profile.
From this control torque evolution, the peak torque value can also be estimated. Moreover, the accumulated angular momentum due to these perturbations are obtained from the integration along the time in which they are acting.
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Solar Radiation Pressure: As this perturbation depends on the distance to the sun, it can be considered that the force is constant and caused a total momentum storage equal to a constant torque multiplied by the total mission time.
Aerodynamic Drag: Constant conservative value has been considered given by the maximum velocity at the lowest altitude. Regarding the accumulated angular momentum, it has been estimated according to the number of low altitudes passes the S/C undergoes during the initial mission phase.
Control Torque: Tailored mainly to the payload pointing requirement, for instance pointing towards the Moon during the fly-by, as well as driven by attitude corrections during thruster operation mode.
The use of 3-axis control is most versatile for frequent reorientations. Hence, it has been discarded spin stabilization or other type of control type as they do not offer the required level of accuracy. For 3-axis stabilization the most common actuator are reaction wheels and thrusters which may require articulated payload. In the following, 3 possible actuators are analysed, and a final decision justified.
DESIGN 1. Magnetorquers DESIGN 2. Reaction wheels (Including desaturation system) DESIGN 3. 6-Thrusters System
DESIGN 1: Magnetorquers
They use electromagnets to generate magnetic dipole moments. The provided torque is proportional and normal to the Earth magnetic field (leading to loss of effectiveness at high orbits). To estimate the electrical torque through the torquer to create a magnetic dipole (D) that results in 푇 a torque (T) in the vehicle the expression 퐷 = is to be used, where the worst case Earth field B is 퐵 estimated to be4.5 ∙ 10−5.
Therefore, magnetorquers can be used to compensate attitude drift but in general they require much more time than thrusters. This solution was discarded due to the fact that the Earth Magnetic Field decreases inversely proportional to 푟3 , being 푟 the distance from the satellite to the Earth. Hence, to obtain the peak momentum imposed by the mission is, by far, not feasible. Therefore this solution is completely discarded.
DESIGN 2: Reaction Wheels
Reaction wheels are devices that exchange angular momentum about the given principal axis. When sizing them, it is important to distinguish between cyclic and secular disturbances, and between angular momentum storage and torque authority.
For 3-axis control systems, cyclic torques build up cyclic angular momentum in reaction wheels, as the wheels provide compensating torques to keep the vehicle from moving. Typically, the angular momentum capacity of a reaction wheel (limited by its saturation speed) is sized to handle the cyclic storage during an orbit without the need for frequent momentum dumping. As before mentioned, a relevant fact is that if the wheels reach their maximum speed (as a result of continuous torque), they need to be unsaturated by an external torquer like thrusters (momentum dumping).
Therefore, the sizing of the rotation wheels is driven by:
- Torque Authority, determined by the peak disturbance torque in order to keep pointing accuracy.
- Angular momentum storage, sized to handle cyclic storage during one/several orbits without need of frequent dumping. It is important to mention that it has been assumed a worst case
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scenario in which all required angular momentum, either due to perturbation corrections or pointing requirement, is accumulated at the three axes. This is a conservative approach since during the mission these torques will be affecting different axes and the angular momentum will be distributed between the three of them.
To define the most restrictive operational conditions in which the ACDS shall guarantee the mission requirements, the overall mission has been divided in three phases according to their different characteristics.
1) Close Earth Phase:
The main idea is to account for the aerodynamic drag besides the Earth gravity gradient and the Earth magnetic field. The aerodynamic drag will produce an extra momentum storage and a maximum peak that the S/C needs to handle.
This phase is comprised from the beginning of the mission at GTO orbit, to an altitude in which the atmosphere density is negligible and aerodynamic drag is therefore no longer acting. For the present mission trajectory this is can be estimated as approximately ten days.
2) Transfer Phase:
Atmospheric drag is now negligible, therefore the satellite only needs to correct gravity gradient, SP and magnetic field perturbations, being able to apply the thrust in the proper direction. This phase covers almost the whole mission duration and it defines the amount of angular momentum change.
The way to evaluate the thrusters correct pointing requirements is explained based on the known S/C state vector, being possible to evaluate the angular velocity and by differentiating it with respect to time, evaluate the required control pointing torque.
푇푐표푛푡푟표푙−푇푟푎푛푠푓푒푟 = 퐼휃̈
Regarding the accumulated angular momentum along the phase, it is gathered by the integration of the control torque law in time.
3) Fly By Trajectory:
Stands for the final phase of the mission. The Earth magnetic field can be neglected as well as the Moon atmosphere drag. So, the attitude control would just correct the moon gravity gradient and guarantee the pointing manoeuvre towards the Moon centre.
The estimation of the control torque for the pointing manoeuvre depends on the fly-by trajectory. Hence, using the relative velocity of the S/C when entering the moon SOI, the altitude at the periselenium, and the radial position evolution with respect to the Moon, the evolution of the angle 휃 with time can be obtained according to the following hyperbolic curve relation:
1 푟(푡) 휃(푡) = acos ( ( − 1)) 푒 푝 FIGURE 10- Fly-By Trajectory Definition
Where푝 = 푎(1 − 푒2), being 푒 the hyperbolic orbit eccentricity and 푎 its semimajor axis. Both orbit parameters are defined according to the relative velocity at the entrance of the Moon SOI and the radial distance to the fly-by orbit periselenium, in particular for the present mission:
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푘푚 푅 = 1738.1 + 100 = 1838.1푘푚 푉 = 0.46 푝 ∞ 푠
Furthermore, taking the 휃 second derivative with respect to time, it is possible to obtain the required angular acceleration which can be converted in a control torque according to the expression:
푇푐표푛푡푟표푙−퐹푙푦퐵푦 = 퐼휃̈
To finish, integrating this torque in time, the accumulated angular momentum during the transfer phase is to be determined.
Some other possible peak torques that may appear during the mission have been considered in order to ensure that the maximum control torque is not exceeded. Those are remarked hereafter.
- Thruster misalignment with respect to the centre of gravity: Which can be caused by an error in the engine installation or be due to the movement of the cg throughout the mission.
푇푡ℎ푟푢푠푡푚𝑖푠푎푙𝑖푔푛푚푒푛푡 = 퐿 · 퐹푡ℎ푟푢푠푡
The worst case scenario has been calculated according to the main propulsion subsystem selected and accounting for a maximum misalignment equal to 20% the transversal distance of the satellite.
- Changing attitude between mission modes: For instance, change from mission transfer correction mode to science operation mode requires a 90° turn in yaw attitude to be performed in a short time compared with the overall mission duration. As shown [2], the slew torque for 90-degrees slews to be performed in 10 minutes can be calculated as:
2 푇푠푙푒푤 = 4 · 휃 · 퐼/푡
Summary of “Design 2” values
Mission Phases Momentum Storage [mNms] Peak Momentum [mNm] Close Earth Phase 990.02 0.005 Transfer Phase 5519.55 0.005 Fly By 6.99 0.003 Other torques Thruster Misalignment - 0.028 Slew Manoeuvre - 0.001
Design Drivers 6516.56 0.028
TABLE 5- Design Drivers
Table 5 gathers the expected requirements that the AOCS must guarantee in order to let the satellite fulfil its mission. The design drivers are to be taken as the most restrictive values along the mission. Based on them, the set of reaction wheels that are to be used in the S/C have been selected.
Obtained results meet expectations: the highest amount of momentum storage is obtained for the transfer phase, which implies that it would drive the amount of desaturations required and the RCS sizing. Moreover, it is observed a low momentum storage during the fly-by, which means that desaturations are not needed, similarly to the close Earth Phase. Lastly, the maximum torque is resulting to be driven by possible thrust misalignments, although it is important to highlight that a very conservative case has been selected. Therefore the larger torque is likely to be tailored during the fly-by, in particular at the periselenium, where the angular acceleration of the S/C is higher and consequently also the torque.
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Furthermore, notice that the thruster pointing requirement appears to impose no remarkable restrictions that affect the subsystem sizing.
Rotating wheels and RCS thrusters selection
The wheels must selected so that they ensure the design maximum torque provision. The number of required desaturations and the time between them are obtained as outputs from their maximum momentum storage capabilities applying the following expression where ℎ푚𝑖푠푠𝑖표푛 is the expected accumulated angular momentum during the entire mission and ℎ푤ℎ푒푒푙 is defined by the maximum momentum storage of the selected rotating wheel, or in other words, the technology to be used.
푁푢푚푏푒푟표푓푑푒푠푎푡푢푟푎푡𝑖표푛푠 = 푛 = ℎ푚𝑖푠푠𝑖표푛/ℎ푤ℎ푒푒푙
With respect to the sizing of the RCS thrusters for desaturation purposes the following steps have been followed:
1) The force level for momentum dumping that the thrusters must apply is defined as: ℎ푤ℎ푒푒푙푠 퐹푡ℎ푟푢푠푡푒푟푠⊥ = 퐿푡푑푒푠푎푡푢푟푎푡𝑖표푛
Where 퐿 is the moment arm, 푡 the burn time for desaturation and 퐹푡ℎ푟푢푠푡푒푟푠⊥ the component of the thrust generating torque. As observed, the higher the arm the lower the required force by the thrusters. Moreover, the higher the force, the lower the time to desaturate the wheels.
2) The amount of propellant mass required to desaturate one wheel is defined according to the next formulas where 퐼 stands for the total impulse to desaturate a wheel, 퐼푠푝 depends on the propellant and 푔 is the Earth gravity constant. Notice that, despite the fact that only 퐹푡ℎ푟푢푠푡푒푟푠⊥ is useful to desaturate, the propellant mass must be estimated for the complete thrust generated by the engine.
퐼 푀푝푟표푝푒푙푙푎푛푡푝푒푟푑푒푠푎푡푢푟푎푡𝑖표푛 = 퐼 = 퐹푡ℎ푟푢푠푡푒푟푠 ∙ 푡푑푒푠푎푡푢푟푎푡𝑖표푛 퐼푠푝∙푔
However wheels need to be desaturated several times, and therefore the total propellant mass comes from multiplying the previous result by the number of desaturations and the number of 퐼 wheels. 푀푝푟표푝푒푙푙푎푛푡푡표푡푎푙 = ⋅ 푛 ∙ #푊ℎ푒푒푙푠 퐼푠푝∙푔
3) To finish, the total mass of the AOCS system, including reaction wheels mass, the thruster mass and the required propellant mass but excluding the attitude sensor is defined as:
푀퐴푂퐶푆 = 푀푝푟표푝푒푙푙푎푛푡푡표푡푎푙 + 푀푡ℎ푟푢푠푡푒푟푠 ∙ #푇ℎ푟푢푠푡푒푟푠 + 푀푤ℎ푒푒푙푠 ∙ #푊ℎ푒푒푙푠
Following this procedure, the several reaction wheels, which are suitable for the amount of perturbation correction and momentum storage have been analysed.
Variable / RW RWP050 RWP100 RWP500 RW1 Momentum Storage [mNms] 50 100 500 1000 Maximum Torque [mNm] 7 7 25 100 Peak Power [W] <1W <1W <6W <9W Dimensions [mm3] 58x58x58 70x70x25 110x110x38 110x110x38 Mass [gr.] 240 350 750 750 # Desaturations 130.3 65.1 10.3 6.5 Time Between Desaturations [days] 3.0 6.0 30.3 60.5
TABLE 6 – Reaction Wheels Catalogue
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This way the most optimum solution can be selected. For any of the above wheels, gathered at Table 6, the peak torque is not the restrictive parameter since all of them generate higher maximum torque than the maximum required, even leaving some margin. Nevertheless, important facts can be extracted from the comparison.
- The smaller the reaction wheel, the higher the number of desaturations although lower time to desaturate. The only point to be considered is that the desaturation time has to be much lower than the expected time between to desaturations, which is defined according to the mission duration. 푡 푡 = 푚푖푠푠푖표푛 푡 ≪ 푡 푏푒푡푤푒푒푛퐷푒푠푎푡푢푟푎푡𝑖표푛푠 푛 푑푒푠푎푡푢푟푎푡𝑖표푛 푏푒푡푤푒푒푛퐷푒푠푎푡푢푟푎푡𝑖표푛푠
- Selecting the smaller reaction wheel the system mass decreases. - The amount of propellant mass does not depend on the RW selected, since it is linked to the mission momentum profile.
According to information from at Table 6, the smallest wheel has been selected (RWP050), since it offers a reduction in mass budget although it is necessary to desaturate a higher number of time and consequently frequently, this does not have detrimental effects on the mission operation provided that 푡푑푒푠푎푡푢푟푎푡𝑖표푛 ≪ 푡푏푒푡푤푒푒푛퐷푒푠푎푡푢푟푎푡𝑖표푛푠. Moreover, the smallest one fits better within the S/C.
Regarding the desaturation system, an option is to use a secondary RCS system. To allow desaturation in the three axes, at least 6 thruster would be required. First, Cold Gas Thrusters have been considered, increasing the subsystem weight up to 3855.4 g. According to the formulation shown previously, it has been assess that increasing the propellant Isp would lead to a decrease in the propellant mass, reason why, a second monopropellant with lower dry mass has been evaluated (hydrazine rocket). However the reduction in mass is mainly driven to the reduction in thruster dry mass so that the S/S can be reduced to 2736.2 g. Furthermore, an ion thruster has also been analysed. The pertinent results from the three different options are gathered in Table 7.
Liquid Rockets 1N Variable Cold Gas MEMS BUSEK BIT-1 Hydrazine Thruster Total Impulse 989000 135000 - Thrust force 53 1000 0.1 Propellant Isp 50 220 2150 Dry Mass 456 290 53 Desaturation Time [sec] 3.77 0.2 2000 Total Mass AOCS System (excluding 3855.4 2736.23 1281.7 sensors) [gr.]
TABLE 7 - RCS Evaluated Systems Information and Results
As it can be observed, for such a light and small satellite it is not worth it to provide a complete RCS system. The level of propellant is negligible for the considered tiny wheels and therefore these options are completely discarded. As most of the weight comes from the thruster dry mass, the final solution is achieved with the main propulsion system, which is turned OFF for the 30% of the transfer time. Therefore, the BIT-3 System Configuration which allows to implement a 2-axis, ±10° gimbal is perfectly suitable to be used to desaturate the small reaction wheels.
Nevertheless some constraints are to be added to the problem:
- The main thruster can only be deflected 10 degrees from the main direction for thrust vectoring. Therefore the force component capable to generate torque is reduced. Despite this reduction, the results are shown to be valid mainly due to the small torque requirements.
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- Though this setup is under-actuated for roll-axis control, the thruster can theoretically offload the angular momentum through a 3-burn manoeuvre. Moreover, a redundant wheel (4th wheel) is proposed to alleviate the RPM requirement on the roll axis so the thruster’s role in angular momentum dumping is minimized. [1]
Variable BUSEK BIT-3 As a result an extra propellant mass of 15g is # Desaturations 130.33 expected to be required for desaturation, Desaturation Time 687.1 which will not have any impact on the Time Between Desaturation 3.0 propellant budget. On the other hand, Propellant mass per wheel 5.11 desaturation will take place every 3 days and Total Mass AOCS System will last 11.5 minutes. 975.33 (excluding sensors)
TABLE 8- Final Solution for Desaturation
DESIGN 3: Thruster system
The sizing of a thruster system as attitude control S/S is quite similar to its sizing for desaturation purposes. However, instead of being designed to counteract the wheel capacity momentum storage it is designed to counteract the disturbance torque multiplied by the time, also called ℎ푚𝑖푠푠𝑖표푛.
Consequently, this design option implies higher amounts of propellant due to the need of 6 thrusters to control attitude along the 3-axis. In case the primary propellant system would not be selected as desaturation system, this solution would have led to better results that including reaction wheels plus a secondary RCS to desaturate. As this is not the case, the following reasons drove the discard of this design option.
- Increased complexity of the ADCS S/S, causing more failure probability. - Higher mass - Need of new tanks and pressurization systems in the case that a different propellant were used instead of the one of the primary propulsion S/S. - The mission time shall be taken into account since the number of impulses that would be required to avoid losing attitude accuracy, will be larger.
SENSOR SELECTION The effect of the accuracy requirement during science operation mode leads towards star tracker or horizon sensors and possibly gyros as selected [2]. Therefore, a sensor analysis has been performed to evaluate which are the more suitable to meet the mission requirements. Notice that sensor technology is evolving rapidly, promising more accurate, lighter-weight sensors for future missions Regardless, data from different sensors has been evaluated and is gathered at Table 9 accounting for advantages, disadvantages and some additional comments coming from the literature and experience from previous missions. According to the performance parameters portrayed at Table 9, three kind of sensors have been automatically discarded.
- GNSS: Nowadays only offers low accuracy, not suitable for requirement AR-007 compliance. - Horizon Sensors: Reach their higher accuracy at low-Earth orbits. Therefore it is not valid for the current mission. - Magnetometer: Is not usable above 6000 km altitudes.
For the remaining options, a trade-off is to be performed. Features as cost, suitability for the current mission, mass and power consumption have been rated and given an importance level. This way, s tar sensors are selected in combination with a gyro. Despite their high cost and mass, together they offer the accuracy that it is searched for the present mission profile.
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TABLE 9 - Sensor Characteristics
During the mission, the number of eclipses will be large at low altitude, however as the distance increases the eclipses diminish. Since electrical propulsion is used and therefore intermediate orbits are covered before reaching the Moon distance, several of them are performed at low altitudes. Then, the sun sensors will not be operative during a high amount of time, since they are visible-light detectors which measure one or two angles between their mounting base and the incident sunlight [2]. For this reason, and because Star Sensor market has increased the efficiency in applicability to nanosatellites, star sensor has been used as a first design solution for attitude control. Nevertheless, the abovementioned set of sun sensors is still required for the sun pointing mode. [28]
This one, represents the most common sensor for high accuracy mission. It derives the attitude of the vehicle after several star crossings. Notice that special attention is required in its specification and use. For instance, the vehicle must be stabilised to some extent before the trackers can determine its position. Also, star sensors are susceptible to be blinded by the Sun, Moon, or even planets. That is the reason why, for high accuracy missions a combination of star trackers and gyros is a common option. In fact, the gyros are used for both initial stabilization and for the periods with Sun or Moon interference in the trackers, while trackers are used to provide the accuracy low frequency external reference unavailable to the gyros.
Nevertheless, the associated complexity and mass of star trackers and the resources that they impose, have limited their use in nanosatellites. Although this is no longer the case due to new advanced design and creative algorithm construction, as the one of the ST-200 star tracker [27], which was created by Hyperion Technologies in cooperation with Berlin Space Technologies GmbH. It is one of the smallest and lightest fully autonomous, low power star tracker, aimed at applications in pico and nanosatellite platforms.
This system features an internal gyro that allows for the determination of the slew rate up to 200 degrees per second even without a visual lock on a star. In order to guarantee correct star tracker operation, the camera position in the spacecraft shall point towards the reference stars.
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Finally, the selected sensors are summarised at Table 10.
Star Sensor + GYRO - ST200 6x NSS CubeSat Sun Sensor Accuracy [arcseconds] / [deg] 30 <0.5 Nominal Power Consumption [mW] / [mA] 650.00 <10 Peak Power Consumption [mW] 1000.00 - Mass [gr.] 42.00 <5 Supply Voltage [V] 3.5-5.0 5.0 Operating Temperature [ºC] -20 to 40 -25 to 50
TABLE 10- Selected Sensors and Back-up system
However, as a back-up system as well as to ensure solar arrays pointing during safe mode six sun sensors were installed in each of the S/C faces. This reduces the mission risk without high impact in mass. Missions as SMART-1 [4] and Lunar IceCube [1] with similar mission goals, use Star trackers for
Science Mode plus Sun sensors for Safe mode/Sun acquisition mode.
CONTROL ALGORITHM A schematic control algorithm is shown at Figure 11. In particular, it is the one used in SMART-1 mission. As it can be observed, from measurements an attitude value is estimated, which is compared with the target attitude. At this points, the control algorithm is in charge of the activation of actuators to control the spacecraft dynamics.
Nevertheless, several algorithm can be found in the literature. Two examples are briefly described in order to present some alternatives that can be used for this topic. [5]
FIGURE 11-Control Algorithm Schematic SMART-1
Active attitude control using a PID algorithm:
A PID algorithm controlling the level of torque is theoretically capable of producing zero attitude error, even when there is a constant disturbance torque. At Figure 12 the response of the roll error and the control torque due to PDI algorithm in the presence of a disturbance torque is provided. FIGURE 12 - PDI algorithm
On/off: Limit Cycle
With on/off control, the limitation of having only three torque levels— clockwise, zero or anticlockwise—means that the control objective becomes the maintenance of an attitude error within acceptable bounds. With on/off control, the error φ is expected to be settled into a limit cycle under steady conditions. FIGURE 13 - Roll Error ΦLimit Cycle
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3.5 COMMUNICATION SUBSYSTEM AND GROUND SEGMENT
The preliminary design of the telecommunications subsystem presented along this section is mainly focused on determining the Link Budget as well as the antenna configuration and the Ground Station Selection.
Link budget depends only on the orbit of the satellite and the main characteristics of the transmitter and receiver. The data budget depends not only on the orbit, but also the methods used to stablish the communication.
SUMMARY OF THE MAIN S/S CONSTRAINTS/ ASSUMPTIONS See Appendix B – Requirements for further details
- A timespan of a week is required to complete the transmission of 10MB of data collected within the Moon fly-by. (See CR-010) - A data rate transmission availability of 264.55 bits/s when communications with Earth are available. - A data link with enough power to transmit the required data packages from a distance equal to the Earth-Moon distance. - The receiving Ground Station shall be within the ESA Network and provide the bandwidth required for the data rate considered. (See MR-009) - Antennas Design shall search for a minimum size and weight, reducing the overall mass of the system. - Location of the antennas shall allow the correct functioning of the rest of the subsystems not covering the solar panels. - A margin of 3 dB shall be accounted for in the link Budget plan. (See CR-019) - The elevation angle shall be selected with the objective of reducing the atmospheric losses. Hence, the value of said losses shall be less than 1dB. (See CR-017). - The maximum and minimum Earth Center Angle, for correct communication link is respectively 0 and 10 deg. (See CR-014) - The telecommunications mode considered shall be Science and Telecommand. Telemetry bitrate will be assumed negligible although existing. - One antenna shall be used for uplink and downlink transmission.
The following frequencies are the ones to be considered for the trade-off analysis (See MR-010): - S-Band with ranges [2 – 4 GHz]. - Ultra-High Frequency (UHF) with range [200-450 MHz].
COMMUNICATIONS CONFIGURATION SELECTION Among the different architectures that can be used to transmit data the following three are to be considered: point-to-point, relay architecture and constellation. The first one consists on sending the information from the S/C directly to the GS. Secondly, relay architecture is usually performed when a nearby S/C has a better position with respect to Earth, thus the information is sent towards this S/C and lastly to the GS. Lastly, the constellation architecture allows communication among satellites. Therefore, the selected architecture in the agreement with the mission design is a point-to-point communication link, saving this way the necessity of designing an additional link among satellites.
Data Rate
The trade-off is to be performed using the two frequencies band abovementioned. The quantity of data that shall be received by the Ground Station for one fly-by is 10MB that is, 80 Mbits (See CR- 001). The time available to perform the transmission of the data without endangering the mission is a week with a nearly continuous data transfer beginning the transfer from an Earth-Moon distance.
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Thus, the data rate can be obtained applying the following expression where 푇𝑖푛𝑖푡 is the initialization time for the connection, which is considered negligible at the present design stage. Notice that the constraints imposed to the S/S are mainly related to the amount of data to be downloaded and that the TTC does not impose restrictions on the S/S since that amount is small.
퐷 푅 = = 264.55푏𝑖푡푠/푠 푇푚푎푥 − 푇𝑖푛𝑖푡
Selection of the Modulation
In order to obtain the link budget, first the modulation shall be selected since this latter one affects the signal-to-noise ratio.
The important aspect of the modulation is the efficiency (spectrum utilization) which is defined as the data capacity that can be transmitted by means of a certain modulation. Considering Quadrature Amplitude Modulation QAM which is very effective as it admits a large number of symbols, the efficiency is determined as the logarithm of the number of symbols defined in the constellation. By selecting 4 symbols, the efficiency obtained is 2 bit/sHz. This scheme, with the four symbols shown at Figure 14, corresponds to the QPSK modulation.
This modulation allows a good performance and an excellent use of the spectrum. Moreover, together with the accounting of an unmanned S/C FIGURE 14-QPSK condition, the modulation leads to a signal- to- noise ratio of 9.6dB. Modulation Symbols
Losses of the transmission link
The following losses are considered for the final link budget calculation, the values for both the UHF and S- Band are going to be stated.
- Transmission Path Losses: The losses from the transmitter to receiver (Lr) are taken into account herein setting a value of 3 dB from reference [2]. Additionally, the transmitter to
antenna loses are taken into account (Ll) by assuming a 5 dB loss.
- Antenna Pointing Loss: Although, the ACDS shall provide the required pointing for the antenna, due to any possible pointing mismatch 1dB is added.
- Thermal Noise losses: These losses are assumed constant and can be obtained with the following expression: 1 푇ℎ푒푟푚푎푙푛표𝑖푠푒 = 푘 ∙ 푇푠
Where 푇푠 is the noise temperature with a value of 500 K for the UHF and S-Band. Therefore, those losses (Lt) get a value of 200dB.
- Atmospheric Losses: The following figure allows to obtain the attenuation at 90º of elevation (zenith) due to the presence of atmospheric gases.
Considering 0.4 GHz frequency for the UHF, 4 GHz frequency for the S-Band, and the change in azimuth, the following expression can be applied in
order to obtain these atmospheric losses (La), which for a 20º of elevation, stand for 0.14dB in the case of the UHF and for 0.10dB for S-Band. FIGURE 15 - Atmospheric Losses
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푑퐵 퐿 = 퐴퐴 ( ) 푐표푠푒푐(휖)푇 (푘푚) 푎 90º 푘푚 푡푟표푝 Space Losses
The space losses are highly influencing due to the great distances considered for the transmission from the Moon back to the Earth, thus they will impact the link budget greatly. These losses depend on the used frequency band, increasing as gathering higher ones. The expression to be considered in order to obtain their value is the following:
퐿푠 = 20 log(3퐸8) − 20 log(4휋) − 20푙표푔푆 − 20푙표푔푓
The obtained losses are 197dB for UHF and 210dB for S-Band. Additionally, to the link budget a margin of 3 dB needs to be added at this early stage of the S/S design.
GROUND STATION SELECTION Based on the requirement of the frequencies available for transmission, as well as on the requirement that is related to the ESA Network Ground Station, two possibilities are taken into account.
For the S-Band range of frequencies the selected Ground Station is KOUROU, since it is known to allow the S-Band and X-Band communication. The antenna available is a parabolic reflector antenna with 15 meters of diameter. Although it is not usually used for this type of missions, typically for early orbit phase, it is still being considered for this mission due to a good accessibility for the communication. Thus, by using the gain formula corresponding to parabolic antennas and by assuming a 0.8 efficiency, the obtained gain gets a value of:
∙ 퐷 2 퐺 = ( ) ∙ 휂 = 45푑퐵 푟
This gain is reduced to 30 dB for practical considerations, mainly because this station does not commonly support this sort of missions involving nanosatellites. Therefore, the power availability is to be reduced.
On the other hand, for the Ultra High frequency range, the selected Ground Station purely follows educational purposes. It is located in Bologna and belongs to the University of Bologna although it does belong to the ESA Network. The available antenna is a Yagi antenna of 2x19 elements providing a gain of 15dB. In Appendix I – Additional Telecommunications s/s information more information on the selected GS can be found.
ANTENNA SELECTION To select the antenna, commercial websites have been consulted with the objective of identifying the antenna that could handle the imposed requirements. The typical gain for S-Band antennas could be in the range of 8-10dB for nanosatellites. Hence, the selected antenna has a gain of 8.3dB. Its evolution along the supported frequencies are provided at the Appendix I. Notice that the frequency selected for the sizing of the power budget is the one that allows the maximum gain.
In the UHF case, the antennas chosen are dipoles that typically have a 1.6dB gain.
LINK BUDGET CALCULATION AND TRADE-OFF The link budget is to be obtained for the most restrictive case, which is the one that accounts for the transmission of data from the Moon distance to Earth. Nevertheless, by varying the elevation angle, it is possible to achieve a lower amount of power required. This one is obtained applying the
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following expression and transforming into decibels the obtained power value. Notice that a power ratio can be transformed into dB by simply using the logarithmic transformation10 ∙ log 푥.
퐸푏 푃 ∙ 퐿푠 ∙ 퐿푎 ∙ 퐿푙 ∙ 퐿푟 ∙ 퐺푟 ∙ 퐺푇 = 푁0 푘 ∙ 푇푠 ∙ 푅
퐸푏 The former equation leads to the expression: 푃 = + 푘 ∙ 푇푠 + 푅 − 퐺푟 − 퐺푇 − 퐿푟 − 퐿푙 − 퐿푎 − 퐿푠 − 퐿푚 푁0
So that the required amount of power obtained for the two configurations is summarized at Table 11 where it can be appreciated that the power value of the UHF is not included is notably larger, showing its lack of capability to transmit the defined amount of data with the used modulation and the considered transmission time. Notice that with the obtained power values, the trade-off is easily assessed since in terms of dimensions and weight, both antennas do not impose high requirements so the selection of the band is correctly tailored based on the obtained power values.
Analysed Band Power Value (W) UHF 155 S-Band 0.6
TABLE 11- Link Budget
Summary of the selection
The selected antenna/frequency band stands for an S-Band patch antenna for S-Band transmission. The specifications of that particular antenna are gathered in the following table:
Specification Value Gain 8.3 dB Half Power beam width 70 deg. Dimensions 98x98x5 mm Temperature range -20 to +60 °C Weight 64 gr. Power required 1.5 W
TABLE 12-Antenna Specifications
Additionally, the telecommunications subsystem requires a transceiver [11] which allows the uplink and downlink transmission. This equipment is selected from the available commercial off-the shelf characterized by the properties gathered hereafter.
Specification Value Dimensions 87x93x17 mm Temperature range -40 to +70 °C Weight 190 g Power required 0.55 W
TABLE 13-transceiver specifications
Free Space Latency
Lastly, to conclude the telecommunications S/S preliminary design, the free space latency is to be determined indicating the delay of the signal across the free space. The large distances resulting from the assumption of Earth-Moon distance for the sizing of the S/S are the most conservative case to be considered so it is assumable that the connection will take place with no hitch. Thus, the latency is computed applying the expression:
푑푐푟𝑖푡𝑖푐푎푙 푡 = = 1.28푠. 푙푎푡푒푛푐푦 푐
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3.6 ELECTRIC POWER SUBSYSTEM
The preliminary design of the Power S/S is divided into three sections. First, a constraint highlight, then a configuration selection and trade-off and lastly, a sizing and summary of the performed selection.
SUMMARY OF THE MAIN S/S CONSTRAINTS See Appendix B – Requirements for further details
- The EPS system must be designed for a 5-year mission. (See MR-008) - All solar panels shall be folded and do not disturb during the launch phase of the mission. The batteries shall be charged before launching. - The Power S/S shall provide enough power to allow the functioning of the required subsystems. The peak power provided shall be 54.15 W in daylight configuration.(See ER-009) - The nominal power available during the transfer for the correct achievement of this manoeuvre shall be 45 W.(See ER-010) - The nominal power available during the fly-by phase where the payload is in functioning mode shall be 20 W. - The batteries shall provide with a 5% of the nominal power, in case of a safe mode operation shall be induced over 12h. (See ER-003) - The loss of power along the transmission lines considered shall be of a value equal to 10% - Two considered mode operations are: First, propulsion at nominal power along with telecommunications in receiving mode and ADCS at nominal power; Secondly, all subsystems at nominal power except the propulsion subsystem that is in standby mode. - A conservative 10% of margin is to be applied to the total power budget (See ER-008) - During eclipse the propulsion subsystem shall be functioning in safe mode operation. The batteries in this case shall provide the power for the rest of the subsystems.
Here, the power required for the nominal and safe mode operation of the each S/S is provided:
Subsystem Nominal Power (W) Peak Power (W) Safe Mode Power (W)
Propulsion 40 40 2 Attitude Control 10 18 2 Thermal 2 2 2 Telecommunications 2 2.2 0.2 Computer control 0.55 0.55 0.55 Payload 5 5 0.01
TABLE 14 - Power Requirements
Regarding the presented values, and due to the obligation of maintaining all supplied equipment with adequate power levels, high values of power necessities are observed so the following two scenarios have been considered.
In the transfer phase, the S/S that shall be forcefully supplied are: the Propulsion S/S at its nominal power (40 W), the telecommunications S/S at safe mode 0.2, the ACDS at nominal power and the thermal S/S at safe mode. This leads to a demand of 52,2 W. Nevertheless, it shall be reduced taking into account that the propulsion S/S would only be active an average of 66% of each of the orbits. Thus, with this value and by adding the 10% margin the total power required results to be 45 W.
The distribution of said power during an arbitrary orbit shall be performed so that the system does not reach the maximum value. Thus, by taking a conservative approach, a possible distribution is provided at Figure 16. The total impulse duration shall be in the percentage range 60-66%.
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FIGURE 16 - Power Distribution along Transfer Orbit
When performing the fly-by, and therefore when recording the data, the Propulsion S/S remains in safe mode whereas the rest of subsystems in normal operation. The total power requirement is then reduced in this case scenario to 24 W. Hence, the former case is not restrictive enough so the dimensioning of the subsystem shall be performed considering the transfer phase.
SELECTION OF THE EPS CONFIGURATION In order to provide the power required (peak configuration) there are several configurations that are to be considered. For this matter, and taking into account the mission geometry and lifetime as well as the power that shall be provided, the following figure can be used to confirm which is the most suitable one.
Although, the power required is lower than the range of the figure, it can be still assumed the location inside the represented circle. As a consequence, the primary energy source for the satellite is going to be a set of solar arrays.
For the dimensioning of the selected primary power source, the first taken assumption is based the solar panels orientation, always towards the Sun. Therefore, the obtained power is always the same but when eclipses are present. Regarding the first selection to be performed, that is the material of the solar cells, for which two possible configurations are taken into account: Silicon and Gallium FIGURE 17 - Primary Power Source Selection Arsenide (GaAs).
Sizing of the Solar Panels To perform the sizing of this subsystem, reference [2] has been be taken into account to guide the main steps. Consequently, the first one is to be the definition of the power requested during daylight and eclipse. These values are presented in Table 15 along with the duration of said phases accounting for the most restrictive case (largest transfer orbit) with the mentioned considerations.
Subsystem Power Daylight (W) Power Eclipse (W) Propulsion 40 2 Attitude Control 10 10 Thermal 2 2 Telecommunications 2 0.2 Computer Control 0.55 0.55 Payload 0 0 Total 54.55 14.75
TABLE 15 - Power Requirement during Daylight and Eclipse
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It is to be noticed that the payload is not considered herein due to the fact that it does not restrict the sizing. Mainly this is consequence of the fact that during the fly-by, when the payload is going to be active, the propulsion S/S would, as before mentioned, not operative.
The times for each of the phases are the following:
Daylight (Length During Orbit) [h] Eclipse (Length During Orbit) [h] 655.4 3.4
TABLE 16 - Daylight and Eclipse Length
The just presented eclipse time has been obtained considering the last transfer orbit. Notice that even more restrictive would have been the GTO orbit analysis, for which the obtained time eclipse is high. Nevertheless, this case is considered as unlikely to happen. Thus, the power that the solar array shall provide is obtained through the following formula:
푃 푇 푃 푇 ( 푒 푒 + 푑 푑) 푋푒 푋푑 푃푠푎 = ∙ 1.1 = 45.2푊 푇푑
The transmission and distribution efficiencies are set to 90%, a typical value for this type of configurations. The 10% left is dissipated as heat and considered in the thermal S/S sizing. The daylight power was obtained as a weighted average of all the components with the purpose of applying a 66% on the power of the thruster, as it is going to be in operational mode just for the said percentage of the time. The power emitted by the Sun at the distance of the Earth is 1367 W/m2, considering that the incident angle is always 0º. Hence, all the panels are perfectly orientated towards the Sun so that the total received power by the cells is the one emitted by the Sun.
2 푃𝑖푛푐𝑖푑푒푛푡 = 1367푊/푚
The power provided by Silicon cells at 30 ºC is equal to 202 W/m2, Silicon cells are quite typical and cost efficient. The GaAs cells on the other hand, provide 400 W/m2 with an efficiency of nearly 30%. To account for the inherent degradation that the cells suffer, the previous value shall be multiplied by an efficiency of 0.83 that includes the following components:
- Design and assembly [0.77-0.9] - Temperature of array [0.8-0.98] - Shadowing of cells [0.8-1] - Inherent Degradation [0.49-0.88]
Hence, the power at the Beginning of Life (BOL) for both configurations can be summarised at Table 17, together with the results obtained for the EOL. Regarding this latter one, for the end of the mission, and again considering a conservative approach, 5 years duration has been set as EOL to determine the level of degradation of the solar arrays. Therefore, assuming a life degradation of a 5 3%, the following expression is to be applied and the EOL power obtained: 푃퐸푂퐿 = 푃퐵푂퐿 ∙ (1 − 0.03)
Cell Power BOL (W/m2) Power EOL (W/m2) Silicon 167 144 GaAs 339 291
TABLE 17 - Power available for both configurations at BOL & EOL
To obtain the area of the solar array that shall be provided for the functioning of the system the power per area unit at the end of life is to be used. The main reason for this is to take a conservative approach which considers that at the end of the mission the solar arrays shall provide the full 푃푠푎 necessary power. Therefore, 퐴푠푎 = . 푃퐸푂퐿
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The results for the proposed configurations lead to an area of 0.123 m2 for the GaAs cells and 0.29 m2 for the Silicon ones. Thus, it is clear that the Gallium Arsenide cells allow a much smaller area and thus are much more recommended.
SUMMARY OF THE SELECTION The solar arrays that are going to be mounted on the nanosatellite are going to be made of Gallium Arsenide (GaAs). The efficiency of the selected cells is about 30%. The dimensions of the commercial panels selected in order to fulfil the requirements are summarized in the next Table 18, as well as their weight and temperature range. Based on that selection, a total number of 16 panels is required.
Furthermore, the power required to deploy the panels is 2W, continuously provided during the time of deployment. Additional equipment that has to be considered as part of the subsystem is related to a bus, both converter and supplier of power, and the on-board computer CPU. The first, is in charge of the transformation of the varying potential and intensity coming from the solar panels to a steady voltage and electric output to be delivered to the loads. The second, is in charge of the command and data handling tasks. Their dimensions, together with other properties are as well gathered at the table. All the following equipment was extracted from references [12] and [8].
Component Dimensions (Mm) Weight (Kg) Temperature Range(ºc) Solar Array: 0.123 m2 8 x 82.6 x 211.5 0.36 -40 to 80 ºC CPU 96 x 90 x 12.4 0.1 -25 to +65 °C Bus Converter 96 x 92x 11 0.07 -20°C to +60 °C
TABLE 18 - Solar Panel and Equipment Specifications
SECONDARY POWER SOURCE SIZING The secondary source of power for the nanosatellite is going to consist on batteries. These will be seized according to reference [2]. In particular, account for the power during eclipses. For this matter, recurring to the Table 15, the batteries shall provide 14.75 W. By taking into account the power margin this value raises up to 15.4 W. This sizing is considered as the conservative approach thus including the possibility of having to provide for 12h a 5% of the nominal required power.
The number of cycles of the battery results 442, considering one cycle per orbit. Thus, as electric propulsion is used, the number of orbits is large. The degradation of the batteries is presented as the depth of discharge DOD (%) and for this case the value is 77% (already considering a 10% margin). This value is obtained from the information provided in reference [2] after choosing Lithium-Ion batteries, mainly because those are the mostly used and a commercial specimen is easier to treat.
Regarding the energy density that allows complying with the power requirement, it can be determined by including in the following formula an energy efficiency of 88% and the Lithium-Ion battery specifications [13]. Specification Value - DOD for the required cycles 70%. Dimensions (Mm) 96 X 92 X 27 - Energy density: 38.5 Wh. Weight (Kg) 0.270 - Efficiency: 90%. Temperature Range(ºC) -20ºc To +60ºc
푃푒푇푒 TABLE 19 - Battery Specification 퐸 = = 76푊 ∙ ℎ %퐷푂퐷 ∙
Two of these battery packs shall be included as the secondary power S/S to cover the required power including the margin left for any type of contingencies. Finally, the summary of the specifications for this additional power source are provided at Table 19.
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3.7 MECHANICAL DESIGN AND STRUCTURE
This section describes the foreseen mechanisms and satellite configurations, as well as provides an overall view of the whole nanosatellite with its global properties and mass budget. The software used for the design of the satellite structure and subsystems configuration enclosed in it, has been IDM-CIC that offers engineers a powerful way to design and validate space mission concepts during pre-phase A studies. Moreover, the aid of viewers as IDM-VIEW and Sketch-Up has been used.
LAUNCHER SELECTION AND LAUNCH PHASE DESCRIPTION Although the launch phase is out of the scope of this PDR, the truth is that having some details about it, is quite important for a rough sizing of the S/C structure. For this reason, some research has been carried in order to select a suitable launcher so that the needed overview of the initial requirements for the structure could be easily identified. In particular, the Soyuz launch System, whose main supplier is the Russian Federal Space Agency, has been the selected one. It is part of the European Space Transportation Union and is operated by Arianespace at the Guiana Space Centre. Some of the principal reasons for choosing it are:
- It has already put in orbit 65 CubeSats in different launch campaigns. - It serves the ISL launch services by ISIS that are prone to be the company in charge for our mission. - It is compatible with the chosen deployment systems and launcher sequences, explained in the following subsections.
The Soyuz mission consists in a three-stage sub-orbital ascent and one Fregat burn leading to the injection into the GTO, desired initial orbit by the present project, with osculating parameters at separation resulting in a ∆V requirement on the satellite’s propulsion system of 1490 m/s. Please, found at the Appendix F – Launch Phase Description
Regarding the deployment system used for the launch, it is relevant to mention that CubeSats are typically launched into space as a containerized payload, inside a deployer, as a manner to reduce launch campaign complexity and associated costs. One of the main characteristics of this sort of S/C. Thus, the best option consists on the 6-Unit DuoPack by ISIS Table 20, due to the fact that the 3U deployer imposes a maximum payload weight of 6kg and our satellite’s weight is well above that value. It can be checked that any of the launch environmental loads (mentioned at the Appendix) overcomes these specifications. Thus, the requirements are prone to be FIGURE 18 - ISIS 6-Unit accomplished. DuoPack
Thermal Operation Quasi Sine Random Shock Load Range Static Load Vibration Vibration 20g-100 Hz; 2000g- 2kHz -35 ºC – 80ºC 15 g 4g 10 -100 Hz 14,1grms (3σ) 6000g- 8kHz
TABLE 20 - ISIS 6-Unit DuoPack Operational Characteristics
S/C design characteristics
- To ensure the separation conditions in spin-up mode, the maximum S/C dynamic unbalance standing for the angle between the S/C longitudinal geometrical axis and the principal roll inertia axis shall be: ε ≤ 1 degrees. - The fundamental frequency in the lateral axis of the S/C cantilevered at the interface must be ≥ 15 Hz. On the other hand, the fundamental frequency in the longitudinal axis of a must be ≥ 35 Hz. (No secondary mode should be lower than the first primary mode).
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- The Soyuz launch system offers standard off-the-shelf adapters at Ø937 mm, Ø1194 mm. Thus the chosen adapter must ensure dimensions within those specifications.
CONFIGURATION TRADE-OFF As a first approach, the major considered constraints for the configurations were the space savings and the low tolerance in the deviation center of gravity of the spacecraft.
First configuration: The main features of this configuration are:
- FEEP propulsion, constituting the most of the satellite’s volume. - Three-axis stabilization by means of four Hyperion RW, including one for redundancy. - Deployable Antenna (Dipole Configuration). UHF communication band. - Deployable Solar Panels on the top, in order to avoid damages caused by radiation provoked by the selected electric thruster.
Besides the abovementioned characteristics, batteries were placed close to the propulsion system whereas the OBC just above it.
FIGURE 19 - First Configuration Views
Notice that as explained in section 3.3, 4 FEEP thrusters were required to accomplish the transfer orbit. Therefore, in order to avoid the creation of additional torques when firing them, additional concerns were accounted. For instance, a possible solution in order to counterbalance the impact of those undesired torques on the structure, could be to orientate the thrusters so that their line of action passes in each firing through the centre of gravity of the S/C. However, an easier alternative could be to simultaneously fire thrusters placed in symmetrical locations, meaning that they would be required to perform out of their normal conditions, not providing thrust at their full capacity but instead half of it (the required thrust provided adding the contributions of the two operating thrusters). Nevertheless, for the sake of simplicity, another configuration has been analysed based on the great advantages found with the final selected Propulsion S/S. Notice that, if this first approach were to be followed, requirement PR-006 related to system TRL would be likely to be unmet since additional process would be required to ensure the just mentioned thrusters needs.
Second configuration:
- ION propulsion: saving volume but increasing the weight of the satellite. - 3-axes stabilization by means of four reaction wheels. - Patched Antenna for S-Band bandwidth communications - Again, deployable Solar Panels on the top, although as in this occasion more power is required (propulsion S/S increased power budget for instance) their area is notoriously increased.
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FIGURE 20 - Second Configuration
Moreover, notice that in this case, the batteries are more separated from the propulsion subsystem, with the purpose of enhancing thermal upper limits prevention of those demanding components. Nevertheless, for the colder scenario, some passive techniques may be required to meet the ranges.
Third configuration:
In this case, some rearrangement has been performed. The principal change in the design accounts for a reshaped iodine tank, which now only covers 1,5U, but in a vertical configuration instead the horizontal one of the former case. Nevertheless, some aspects have to be taken into account. The fuel evaporating system is customized for the former configuration, as it is the one already with TRL greater than 6. Therefore, the evaporating system may not work properly at its full capacity with this tank changed shape and it has to be considered as an important risk of the mission. Another consideration may be the thermal consequences of this new distribution of the propulsion elements, including the power generator board and the control valves.
FIGURE 21 - Third Configuration (Science Mode)
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This way, the volume can be reduced to only 3 units, making a compact satellite. It also affects positively to the AOCS, as the mass is lower and so are the inertia moments. It can be achieved a more centred position of the cg as well.
After evaluating the mentioned aspects and trading-off with the different subsystems, taking a great importance the propulsion system, the chosen configuration has been the third one, as the centre of gravity is just slightly deviated from the centre of the cross-sectional area of the S/C and results in the low volume of 3U. Finally, regarding the solar panels deployment, notice the fact that the panels are 360 degrees orientable complying with the requirements of not shading the FOV of the camera and the ensuring Sun pointing feasibility.
For the sake of completeness, a trade-off table proving detailed information of the weighted factor used for the trade-off analysis between the configurations is included at the Appendices, together with resto of S/S tables.
Body axes definition
Sketch-up software has been used to represent the body axes of the satellite in order to situate every subsystem of the model, represented in Figure 22, where the axes can be defined as:
- X-axis: pointing, during the fly-by data discharge, towards the Earth direction. With the sake of easing communications, the antenna is placed over the positive x surface. - Y-axis: perpendicular to the x-axis, undefined direction. - Z-axis: tangent to the orbit trajectory, thus pointing in the velocity direction. Aligned with the thrusters and the main payload which are placed in opposite S/C ends. FIGURE 22 – Subsystems/Components Distribution
As it can be seen, the different components defined in the body axes from bottom to top are listed in the following table:
Unit Subsystem Component PROPULSION - Busek Iodine BIT-3 RF Ion Propulsion System [21]. - AC/DC bus U1 OBC - ISIS Computer with Daughter Board[8] - Wiring system - Auriga Star-Tracker[9] ADCS - Hyperion Tech - HT RW210.15 U2 POWER - Gomspace -NanoPower BP4 F Option[12] - ISIS UHF/VHF Transceiver[10] COMMUNICATIONS - Cycle Space-CPUT S-Band Antenna[9] POWER - 1U Generic Deployable Solar Panel[12] U3 - Gomspace – NanoCam[14]Cover PAYLOAD - Insulator Pad
TABLE 21 – Components Flow-down
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MISSION PHASES CONFIGURATIONS Three main states of the satellite are going to be presented, each corresponding with one satellite action or purpose. - Launch: During the launch sequence, being the most dangerous phase in terms of vibrations and loads that the structure must withstand, all the appendages are folded and the camera is kept and protected by the superior cover of the satellite. Moreover, this is the phase in which the structure has to carry the maximum weight of the fuel tanks, that are set as a 100% of capacity. - Transfer: in this configuration, the deployable solar panels are oriented normal to the Sun in order to exploit the illumination radiation power at its maximum extent.
FIGURE 23 - Transfer Mode Startracker
Antenna
FIGURE 24 - Launch Mode
- Science: The cover is opened and the camera’s articulation activated to start with the capturing of data by the payload over the Moon’s surface. Appendages can be deployed or not depending on the radiation direction and the power needs. For the tanks, 60% capacity is considered when the spacecraft performs the flyby due to the fact that most of the propellant remains unburned as gathered at the propulsion section.
MASS BUDGET The mass budget has been estimated by gathering the preliminary predictions of each subsystem engineer and the different components have been distributed within the modules complying with the performance requirements, but also following a realistic and reasonable assembly sequence. Please refer to Appendix H – Configurations mass Budget where the mass budget and internal volumes are specified for each station segment. Different uncertainty mass factors have been applied corresponding to diverse levels of design development of each component, being for the Preliminary Design 15% the recommended one.
Nevertheless, at Table 22 is provided an excerpt of the system final mass results as an exemplification. Its corresponding complete table found at Appendices gathers the mass budget decomposition for the Launch configuration, which is considered to be the most demanding and the one carrying the fuel tank at its 100% of capacity. With respect to the launch sequence, the sum of the maximum wet masses here calculated was provided to the Propulsion engineer.
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Configuration : cfg.LAUNCH Configuration LAUNCH
▼ NANOSAT Target wet mass [Kg] : + Unit Without margin [Kg] Margin [%] Margin [Kg] Including margin [Kg] % of total Subsystem - Name Quantity Mass [Kg] Margin [%] ▼ Subsystem PROP 1,70 0,00% 0,00 1,70 27,11% ▼ Ion Thruster system 1 1,66 0,00% 1,66 0,00% 0,00 1,66 | BIT3-with 2ax gimbals 1 1,44 0,00% 1,44 0,00% 0,00 1,44 | Prop control valve 1 0,11 0,00% 0,11 0,00% 0,00 0,11 | Tank pressure transducer 1 0,02 0,00% 0,02 0,00% 0,00 0,02 | Iodine tank 1 0,09 0,00% 0,09 0,00% 0,00 0,09 Power generator board+PPU 1 0,05 0,00% 0,05 0,00% 0,00 0,05 ▼ Subsystem PAY 1,00 5,48% 0,05 1,05 16,82% Gomspace - NanoCam 35mm 1 0,97 5,00% 0,97 5,00% 0,05 1,02 Cover 1 0,03 20,00% 0,03 20,00% 0,01 0,04 Insulator pad 1 0,00 20,00% 0,00 20,00% 0,00 0,00 ▼ Subsystem STR 0,54 20,00% 0,11 0,65 10,35% ▼ Tutorial Custom 3U 1 0,54 20,00% 0,54 20,00% 0,11 0,65 ▼ Subsystem PWR 0,99 5,00% 0,05 1,04 16,58% Gomspace - NanoPower BP4 F Option1 0,27 5,00% 0,27 5,00% 0,01 0,28 ▼ Solar Array 2 0,36 5,00% 0,72 5,00% 0,04 0,76 | ▼ Assembly 13 1 0,36 5,00% 0,36 5,00% 0,02 0,38 | | ▼ 2U Generic Deployable Solar 2Panel 0,18 5,00% 0,36 5,00% 0,02 0,38 | | | 2U Generic Solar Panel 2 0,09 5,00% 0,18 5,00% 0,01 0,19 | Equipment 15 1 0,00 5,00% 0,00 5,00% 0,00 0,00 ▼ Subsystem ADCS 1,17 15,00% 0,18 1,35 21,45% Hyperion Tech - HT RW210.15 4 0,24 15,00% 0,96 15,00% 0,14 1,10 Auriga Star Tracker Sodern 1 0,21 15,00% 0,21 15,00% 0,03 0,24 ▼ Subsystem OBC 0,27 15,00% 0,04 0,31 4,95% ISIS Computer with Daughter Board 1 0,27 15,00% 0,27 15,00% 0,04 0,31 ▼ Subsystem COMM 0,13 0,00% 0,00 0,13 1,99% ISIS UHF/VHF Transceiver 1 0,08 0,00% 0,08 0,00% 0,00 0,08 ClydeSpace - CPUT S-Band Patch Antenna1 0,05 0,00% 0,05 0,00% 0,00 0,05 ▼ Subsystem THERMAL 0,04 20,00% 0,01 0,05 0,75% Radiator + Al-cover 1 0,04 20,00% 0,04 20,00% 0,01 0,05 Total dry mass without system margin 5,84 7,48% 0,44 6,27 System margin 0,00% 0,00 6,27 Propellant mass 2,73 15,00% 0,41 3,14 Total wet mass including all margins 9,41 SuperCollapser:End System Without margin [Kg] Margin [%] Margin [Kg] Including margin [Kg] Total dry mass without system margins 5,84 0,44 6,27 Total dry mass including system margins 6,27 Total propellant mass 2,73 15,00% 0,41 3,14 Total wet mass including all margins 9,41
TABLE 22 - Launch Configuration Mass Budget
It is important to remark that, in the mass budget that is shown in Appendix H – Configurations mass Budget for the different configurations, the mass without margin refers to the one that the structures engineer received as an input from the different subsystems. This way, Propulsion, Power and Communications subsystems calculations included their respective margins so that the margin applied in the table corresponds to the rest of subsystems, meaning ADCS, OBC, Structure and Payload Subsystems.
Regarding the Van Allen belts radiation shielding, it is directly included the Al mass of the structure protection cover.
CENTER OF MASS AND INERTIA TENSOR The heaviest component of the satellite is the Iodine tank, placed inside the called U1 and making the center of gravity deviate in z axis from the geometrical center of the structure. Thus, as the propellant content in the tank decreases (configurations transfer and science), it is observable that the cg. deviation decreases.
The inertia axes are affected as well by the subsystems disposition and the deployment of the appendages. Rotational inertia also depends on mass distribution and varies with the axis of revolution selected as a reference. These two quantities are provided in the following tables for each of the mentioned configurations.
Cog Including Mass Margins Inertia Matrix at Cog Including Mass Margins x [mm] y [mm] z [mm] lxx[kg.m²] lxy [kg.m²] lxz[kg.m²] lyy[kg.m²] lyz [kg.m²] lzz [kg.m²] 1,4150 6,6820 144,0452 0,0811 -0,0005 -0,0018 0,0778 -0,0076 0,0100 0 0 65 0,0071 -0,0000 0,0009 0,0072 0,000 0,0001 -6,2416 -2,2736 112,6229 0,0944 -0,0013 -0,0048 0,0908 -0,0111 0,0116
TABLE 23 - Launch Configuration Cg & Inertia
COG Including Mass Margins Inertia Matrix at COG Including Mass Margins x [mm] y [mm] z [mm] lxx[kg.m²] lxy [kg.m²] lxz[kg.m²] lyy[kg.m²] lyz[kg.m²] lzz[kg.m²] 1,4150 6,6820 157,3418 0,16127 0,0079 -0,0012 0,0973 -0,0065 0,0755 0 0 65 0,0091 -0,0000 0,0006 0,0091 0,0000 0,0000 -4,5049 -0,2506 129,4730 0,1759 0,0072 -0,0041 0,1117 -0,0099 0,0768
TABLE 24 - Transfer Configuration Cg & Inertia
Cog Including Mass Margins Inertia Matrix at Cog Including Mass Margins x [mm] y [mm] z [mm] lxx[kg.m²] lxy [kg.m²] lxz[kg.m²] lyy[kg.m²] lyz [kg.m²] lzz [kg.m²] 1,1232 6,6820 161,9033 0,1674 0,0079 -0,0007 0,1035 -0,0075 0,0754 0 0 65 0,0095 0,0000 0,0005 0,0095 -0,0000 0,0000 -4,0613 0,5273 136,054 0,1820 0,0073 -0,0035 0,1180 -0,0107 0,0767
TABLE 25 - Science Configuration Cg & Inertia
NATURAL FREQUENCIES In this preliminary design, the structure is going to be considered as a cantilever beam with respect to the launcher adapter interface. Following the literature guidelines, the equations that are used to
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model the satellite behavior and its first natural frequencies are different in case the applied force is lateral or axial during launch, and both are required to be higher than those aforementioned from the launcher envelope. 3 푓푛푙푎푡 = 0.56√(퐸퐼/푀퐿 ) = 731 Hz 푓푛푎푥𝑖푎푙 = 0.25√(퐴퐸/푀퐿) = 3760 Hz
Table 26 gathers the data for the Launch configuration and Table 27 the obtained results for solar panels during their operation in the other two configurations.
E (Gpa) I(m4) M(kg) L (m) A(m2) 71 8,33333E-06 9,41 0,333 0,01
TABLE 26: Data for Frequencies Calculations
The fundamental frequencies of deployed appendages are modeled as a point mass connected to a hinge whose stiffness (k) can be approximated 10000Nm/rad and where J is the moment of inertia of the appendage component with respect to the hinge which can be estimated for each panel (being b the dimension at which the hinge is attached and t the thickness) as hereafter shown: 1 k 푏3ℎ 푓 = √ 퐽 = 푛sp 2휋 J 12 Transfer Science 푱(풎ퟒ) 4,76e-08 1,06e-05
풇풏퐬퐩(푯풛) 6040 493
TABLE 27: Appendages Natural Frequencies
It can be concluded that all of the calculated frequencies are well superior to the lowest allowed S/C natural frequencies for launch configuration, therefore fulfilling the requirement MeR-005.
MECHANISMS Solar panels deployment
The primary mechanism involves a two-spring system in one axis with a travel limit. The deployable board is fixed to an Al hinge by three bolts. This hinge is mounted in the lateral shear plate of the structure by a system composed by two lugs and shaft. The lugs are mechanized in an enlarged area of that Al shear panel. At both ends of the shaft, two steel springs are located. Those springs are the actuators of the rotation movement. There are two mechanical limits which physically stops the opening rotation: the first one fixed to the shear plate ensuring the correct maximum angle and the second one is a steel flat spring which fixes the panel in the final position reducing vibration problems. Regarding the attachment to the structure, the simplest method is a knuckle joint attached to a boom located at the root of the appendage.
FIGURE 25 - Detail of XATCOBEO SA Deployment Mechanism [17]
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Camera articulation
Riel system articulation pushed by a spring on the bottom of the support pad. The cone hinders the debris to damage the lens and allows for the required FOV to accomplish the mission objective.
FIGURE 26 - Detail of Camera Riel System Articulation.
STRUCTURAL MATERIALS Aluminium 7075-T6: Used for the truss structure, skins and face sheets designed with several mounting locations for components in an attempt to offer configuration flexibility. Widely employed in highly stressed aerospace structures because it provides high strength and low density. Also a 10mm shield for Van Allen radiation protection is added following [2].
Thermal treatment T6 hinders brittleness and increases the strength of the alloy by as much as 30%. Important features are its weldability, workability and availability. Thus, the cover and screen of the camera are also made of this alloy.
Steel (High Strength) Employed in the mechanisms (bolts and threaded parts, main sustainers), due to its high strength and stiffness. It can withstand the high loads caused by the deployment torques during operation, which compensates its higher cost and density.
CFRP: Employed in the Slit-Tube Graphite Composite Booms, which supports the solar array turning and union with the structure. Graphite composites are widely used for lightweight structures that need to carry extremely high loads, manufactured by molding process.
Property Aluminum Steel CFRP Density (𝝆) [KG/M3] 2700 7700 1578 Young Modulus (E) [GPA] 71 200 282 Yield Stress (𝝈퐘) [MPA] 503 1034 586 Ultimate Stress (𝝈풖) [MPA] 572 400 -
TABLE 28- Structural Materials Properties
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3.8 THERMAL CONTROL SUBSYSTEM
The thermal control consists on ensuring a proper temperature distribution over the equipment and structure. Electronics, mechanisms and other devices must be operated within a narrow temperature range. Moreover, some of them have an optimal operating temperature, outside that temperature their performance drops.
The spacecraft thermal control deal with the following heat sources:
- Internal heat generated by components, such rocket motors or batteries. - Heat from the environment: The main source is the solar irradiance at the Earth distance from 2 the Sun. It can be estimated from 푆푒 = 1366푊/푚 , which is the solar irradiance, and accounting 6 for the Earth-Sun orbit semi major axis 푎푒 = 150 · 10 푘푚. - Planetary albedo plays also an important role. This is due to the fact that planets are not perfect absorbers, and reflect part of the solar energy. The planetary albedo can be expressed as follows, noticing that 푎 stands for the bond albedo coefficient, and 퐹 the visibility factor :
푆푎푙푏푒푑표 = 푎 · 푆(푟) · 퐹
- To finish, planets generate temperature internally and therefore they irradiate, approximately as a black body.
Space is not a common thermal environment. The satellite is orbiting in vacuum, hence there is not outside convection or conduction. This means that radiation is the only mean of heat exchange with the environment. Inside the satellite, conduction and radiation between the different parts are the relevant mechanism, although convection may also appear if fluids are used, which is not the case.
The way to control the thermal behaviour can be tailored in different manners:
- Passive Control: Paints, coatings, insulations etc. which operate without any power input or command to operate - Active Control: Devices that consume power to operate and/or need to be commanded. They usually have a higher performance than the former.
The preliminary control design for this project is based on passive techniques using different coatings. The reason is to reduce the consumption of power and therefore optimize the mass, as the size of batteries and solar arrays decrease. In particular, the paint used over the components will be selected in such a way all components work in its range temperature. This design also reduces the level of impact on other S/S. For instance, installing a radiator would require to modify the power budget, as well as the mass of the whole system, which would lead to an iterative redefinition of all the subsystems.
Continuing with the previous idea, changing the painting leads to a different behaviour in terms of emissivity (휀 ) and absorptivity (훼 ) of heat radiation. Any body with temperature radiates to its environment and absorbs all or part of the incoming radiation. The emissivity/absorptivity reflect the difference in radiation emitting/absorption with respect to a black body.
Nevertheless, these properties depend on the spectrum of the incoming/emitted radiation spectrum. As for this mission they are very different because the two radiation-interacting bodies (e.g the Sun and the S/C) are at a very different temperature, the integral absorptivity and emissivity can be quite different.
휆=∞ 휆=∞ ∫휆=0 푃푠푢푛(푇푠푢푛, 휆)훼(휆) ∫휆=0 푃푠푎푡푒푙푙𝑖푡푒(푇푠푎푡푒푙푙𝑖푡푒, 휆)휀(휆) 훼 = 4 푑휆휀 = 4 푑휆 퐵푇푠푢푛 퐵푇푠푎푡푒푙푙𝑖푡푒
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This means that a given material can be a good emitter and absorber at IR wavelengths but a bad emitter/absorber at higher frequencies. Exploiting this fact, it is possible to take advantage of it and select:
- White paint, which has a very low absorptivity for the solar spectrum and good emissivity at moderate temperature (around 300 K). Hence it is a good option to avoid solar radiation absorption and for radiating power away. It is called cold surface. - Black-anodising paint, on the contrary, presents a high absorptivity ( 훼 = 0.88 ) for the solar spectrum and high emissivity (휀 = 0.88) for radiating power again. Therefore, is used to maintain the temperature constant. - Gold plated coating, presents the highest absorptivity (훼 = 0.25) and lower emissivity (휀 = 0.04) of the three considered coatings, being used on those surfaces requiring a higher temperature.
The tool used to carry out the before mentioned thermal control strategy is the commercial software ESATAN - TCM. This software allows to generate the satellite geometry, model the heat fluxes due to conduction, convection and radiation and define the heat loads for different satellite operational conditions assessing therefore different radiative scenarios for distinct orbital positions.
The following assumptions have been taken for the ESATAN model definition:
1) The internal heat generated by the components has been estimated according to the device efficiency. All the power that it is not used for efficient purposes it is then transformed into heat.
푃푑𝑖푠푠𝑖푝푎푡푒푑 = 푃푐표푛푠푢푚푝푡𝑖표푛 · (1 − 퐸푓푓𝑖푐𝑖푒푛푐푦)
Table 29 gathers the resulting power dissipation for every satellite equipment/subsystem. As a summary of the considerations taken into account, it is to be highlighted that: the PPU was considered as a type of battery. The power dissipated by the thruster has been estimated to 10W, which is approximately 25% of the power consumed during an orbit operation.
Component Heat Dissipated [W] Batteries 4.56 PPU Thruster 2.5 CPU 0.36 Telecommunications 0.1 Payload 5 Star Tracker 0.1 Thruster 10
TABLE 29 – Heat Dissipation Definition
2) As a first approximation, a set of materials have been selected for the different satellite components. This way their thermal properties are fully defined.
Material Density[kg/m3] Conductivity [W/mK] Specific Heat [J/kgK] Aluminium 2700 205.0 910 Iodine 4900 0.447 429 PVDF 1780 0.2 1120 Galium Arsenide 5317 55 330 Battery Material 250 205 1000
TABLE 30 – Thermal Materials Definition
3) The types of coating used are specified in Table 31. The payload camera and the solar cell coating are defined according to the properties of the materials used for this type of elements. On the other hand, black coating is used for all the internal subsystems and equipment, except for the
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PPU and the thruster which are covered in white. The antenna is painted in gold so that its temperature could be above its lower limit. External and inner surfaces of the structure are covered in black-anodizing coating since, after several iterations, this was the only option that allowed the requirements to be fulfilled in both orbital scenarios explained below.
Coating Absorptivity Emissivity Black Anodizing 0.88 0.88 Gold 0.25 0.04 White 0.14 0.9 Solar Cell Coating 0.4 0.45 Payload Camera 0.3 0.6
TABLE 31 – Coating Properties
In regard to the scenarios, both considered an elliptical orbit similar to the last orbital pass of the transfer orbit, although with different inclination so as to represent the hottest and the coolest orbit possibilities.
For the sun light scenario it has been considered through an elliptical orbit, whose inclination makes the S/C receive Sun light along the whole orbit, as shown at Figure 41 in Appendix J – Additional information of the Thermal S/S.
For the eclipse scenario, gathered at Figure 40 in Appendix J – Additional information of the Thermal S/S, it has been defined as the case in which Sun-Earth-Satellite are aligned and the eclipse takes place at the perigee of an elliptical orbit. In reality the worst case scenario identifies the eclipse to take place at the apogee, but it would lead to extremely low temperatures, under which the temperature requirements of the components were not fulfilled (See Figure 42 in Appendix J – Additional information of the Thermal S/S.) However, that condition is very restrictive and unlikely to occur since it could be avoided by setting an appropriate orbit inclination. That is, that case has been assumed as non-practical and directly discarded. Hence, it has been assumed as non-practical and directly discarded. Notice that the S/C is expected to have the highest temperature gradient at this condition.
In order to dote the analysis with a greater dose of realism, not all the components have been considered to be turned ON at the same time along both of the studied scenarios. In fact, some cannot be operating according to mission definition and requirements. Therefore, the assumed ones are gathered at Table 32, which shows the working status of the different devices at both scenarios. Notice that this selection maximises the power dissipated in the Sun Light scenario, and minimises it in the Eclipse scenario so that they are both as conservative as possible to this regard.
Component Eclipse Sun Light Batteries ON OFF PPU Thruster OFF ON CPU ON ON Telecom ON ON Payload OFF OFF Star Tracker ON ON Thruster OFF ON
TABLE 32 – OFF-ON Status at both scenarios
An iterative process has been performed trying to apply different coatings to the different elements and ensuring they operate for these two extreme worst case scenarios within their specified range. During the process, the most restrictive elements and key drivers from the thermal control definition standpoint have been the thruster in the Sun Light scenario, and the antenna and the star tracker in
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both scenarios. As it was expected, the minimum temperature took place at the eclipse but with the use of black anodizing coating in most of the subsystems, the power generated by the devices in operation allowed the operational temperatures ranges of the studied thermal nodes to be guaranteed. In fact, all the thermal requirements gathered at Appendix B – Requirements are met.
Figure 27 shows the time evolution of the average temperature in the different elements during the eclipse and sun light orbital cases. Notice that only the most relevant elements have been highlighted. For a complete a clearer view of all 14 components modelled, please refer to Figure 43 and Figure 45 in Appendix J – Additional information of the Thermal S/S.
FIGURE 27 - Components Temperature Evolution for the Eclipse (Left) and Sun Light (Right) Scenarios
First of all, it can be observed how the temperature evolves on the eclipse scenario while it is kept constant in the sun light scenario. In the former, the most critical element is the structure which is kept at -13ºC, although it is still in the allowed range. For the Sun light case, the most restrictive element is the thruster given the large amount of power radiated as it is operative. Nevertheless, its temperature (109ºC), although is high, it is inside the range as well.
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4. RISK ANALYSIS AND MITIGATION
Once the main subsystems which define the space segment of the present mission have been presented, it is of high relevance to analyse the level of uncertainty and risk associated to each of them. It is a crucial step since based on their properly identification and evaluation, the type of response that is to be implemented can be tailored for the sake of mission goals accomplishment. These responses could be avoidance of the risk, transference of responsibility, mitigation of the effects under an acceptable threshold of impact and lastly acceptance.
The performed risk classification is based on the definition of two parameters: the extent of the impact their effect may cause to the overall mission and their probability of occurrence. Both are balanced according to the ranges presented at the following tables, where it can be appreciated that five levels have been used to establish the severity of each of the risks. Notice that the impact is given as a %. This results from the engineers’ judgement relating variables such as cost, level of mission requirements compliance, time, mass and performance among others.
Risk Classification Impact [%] Risk Classification Probability [%] Very Low 5 Remote 10 Low 10 Unlikely 30 Medium 20 Probable 50 High 40 Likely 70 Very High 80 Nearly cert. 90
TABLE 33 - Probability Impact Risk Scoring
Having said this, the Impact-Probability matrix is to be obtained and those risks presenting high severity values treated accordingly. Notice that severity has been defined as the product of the two parameters associated to each of the gathered risks. Therefore, again 5 levels can be identified:
- High severity (Level 4 and 5): P-I score > 0.14 - Medium severity (Level 2 and 3): 0.04 In the seek of contents abbreviation, this section gathers only the analysis and possible solution strategies for the most critical risks. Nevertheless, the entire list of the considered probable risks is provided at Appendix E – Risk Analysis/ Mitigation Table, where the acceptance criteria and mitigation strategies are included. Therefore, based on the criticality grading of the evaluated risks, the following ones are found to be of great relevance: - Failure of the payload: some inoperability situation regarding the camera deployable mechanism may provoke the loss of the main mission objective, making it unworthy. It has to be avoided by means of a severe test and verification campaign of the component. - Main thruster failure: it implies not only the loss of the propulsion S/S itself but the loss of the desaturation system as well. It could be due either to an incorrect input power or mechanism failure. As it is not possible to mitigate, the followed strategy is reduced to the acceptance. Related for instance to the ADCS S/S, the acceptance would shorten the reaction wheels effectivity, implying trajectory deviations and impossibility to orientate the S/C for power inputs. - Low propellant efficiency: a larger iodine tank than the needed one may affect the sublimation process leading to a lower propellant efficiency and, in the worst case, the satellite could not reach its fly-by starting point. This problem must be avoided by the performance of tests for different dimensions and position of the sublimation thermistor. - Damage during launch: caused by the coupling of natural frequencies, being lower than the calculated in the design phase and leading to changes in inertia and cg. It can be mitigated by setting smaller cg deviation margins, in order to avoid severe components destruction or damage if this failure takes place. Project funded by the Interreg Sudoe Programme through the European Regional Development Fund (ERDF) 49 5. CONCLUSIONS AND FUTURE WORK GUIDELINES The design solution presented along this PDR in response to the NANOSTAR Student Challenge Moon fly-by CubeSat, is mainly characterized by the compactness of the satellite. In fact accounting just for a volume of 3U and a mass of 8.84 kg, the nanosatellite has been found to conform a feasible Space Segment in terms of S/S requirements accomplishment. Consequently, high level (Mission requirements) are likely to be meet as well. In fact, although not detailed simulations are portrayed at this phase, the preliminary ones show compliance with them so no deviations are expected. The particular proposed design choices main features, are to be summarized as follows: - Mission analysis: is highly dependent on fly-by specifications, not available at this early stage, so just order of magnitudes are obtained to enable the sizing the S/C subsystems, mainly in terms of worst case scenarios for instance in eclipse definitions. - Space Propulsion S/S: an electric ion thruster model has been selected implying a penalization in mission duration, (18.4 months) in comparison with a chemical approach. Moreover it leads to larger exposure to Van Allen Belts charged particles which affect mass budget by the inclusion of appropriate shielding. Nevertheless the presented advantages such as the available dry mass left to other S/S, or its capabilities to be used as desaturation system with a 2-axis system have been considered to be more relevant. - Attitude Determination and Control S/S: a 3 axis stabilization has been found to be suitable to face the disturbance environment and the required control torques to ensure the correct performance of the S/C in all its operations mode. The selected design choice complies with redundancy and stands for four lower mass RW as the main actuators, which provide an adequate number of desaturations along the mission. The S/S is closed by the selection of a startraker + gyro system as primary set of sensors, and a buck up system of sun sensors. - Telecommunications: the selection of QPSK modulation along with S-Band frequency range enables the downlink of the acquired data along the fly-by into the Kourou Ground Station in a feasible time frame of two days, implying 2W of link budget. Nevertheless, the disposal strategy has been selected to enable the satellite patched antenna to point towards the Earth for 7 days, ensuring the complete downlink of the data. - Electric power: a system of 16 solar panels comprised of GaAs cells is designed as primary power source supplemented with a set of two batteries for proper S/S operation under eclipse condition. The main load has been found to be the propulsion S/S, hence limitations on power budget are set to it, reason why a lower propellant utilization efficiency is used. - S/C Configuration, Mechanisms and Structure S/S: significant level of risk has been assumed in order to optimise the S/C volume, mainly related to the proposed rearrangement of the electric thruster components which will require a TRL demonstration. Nevertheless, requirements are fully met. The designed modular structure is capable of withstanding launch vibrations, it provides a suitable position of cg and inertias values, achieved by a correct disposition of elements. PDR margins have been included to the mass budget and as Al shielding has been applied in the model, it is not included as major risk. - Thermal control S/S: both internal and external heat sources are controlled by means of passive control techniques. In fact, it can be concluded that all the components are maintained within their temperature ranges even in the worst case scenarios which could be avoided by means of a refined mission analysis. To finalise, some of the most relevant future work guidelines are to be highlighted. For instance, in regard to the mission analysis, a proper launch window and targeting analysis which can help to completely determine the orbit are pending for the Detailed Design Phase of the project. Moreover, it would be advisable to investigate other locations along the orbit to perform the thruster firings so that the transfer and the fly-by trajectories allow the possibility to perform a second fly-by. Furthermore, with respect to the propulsion subsystem, an increase in the propellant efficiency becomes necessary, either by investigating the possibility of increasing the power or further researching the propellant tank dimensioning and sublimation system for a performance increase. Project funded by the Interreg Sudoe Programme through the European Regional Development Fund (ERDF) 50 6. REFERENCES [1] Tsay, M., et al. 2015. “LunarCube: A Deep Space 6U CubeSat with Mission Enabling Ion Propulsion Technology”, Proceedings of the AIAA/USU Conference on Small Satellites, Advanced Technologies III, SSC15-XI-1. https://digitalcommons.usu.edu/cgi/viewcontent.cgi?filename=0&article=3235&context=sm allsat&type=additional [2] Larson, W. J. and Werzt, J.R., “Space mission Analysis and Design”, 3er ed. Kluwer Academic, 1991. (2005) [3] Scudder, J., 2017, “Why Aren't The Van Allen Belts A Barrier To Spaceflight?”, Forbes. [Online] Available from: https://www.forbes.com/sites/jillianscudder/2017/06/16/astroquizzical-van- allen-belts-barrier-spaceflight/ [Accessed 07 May 2019] [4] Alonso, M. et al. “SMART-1 Lunar mission: Operational experience with its automatic attitude and orbit control subsystem and its relation with electric propulsion system”, Proceedings of the 6th International ESA Conference on Guidance, Navigation and Control Systems, Loutraki, Greece, October 2005, ESA SP-606. [5] Fortescue, p., et al, “Spacecraft Systems Engineering”, 4th ed. London: Wiley, 2011. [6] E. C. for Space Standardization, “Space project management: Project planning and implementation,” in ECSS-M-ST-10C Rev. 1, 2009. [7] Innovative Solutions In Space B.V., (2016) “CubeSat Deployers” [Brochure]. Retrieved from https://www.isispace.nl/wp-content/uploads/2016/02/CubeSat-deployers-Brochure-web- compressed.pdf [8] Solutions In Space B.V., (2016) “ISIS On board computer” [Brochure]. Retrieved from: https://www.isispace.nl/wp-content/uploads/2016/02/ISIS-On-board-Computer-Brochure- v2R-compressed.pdf [9] Sodern ArianeGroup, (2017) “Auriga” [Brochure]. Retrieved from: http://www.sodern.com/website/docs_wsw/RUB_315/tile_685/AURIGA_Datasheet.pdf [10] Solutions In Space B.V., (2016) “Communication Systems” [Brochure]. Retrieved from: https://www.isispace.nl/wp-content/uploads/2016/02/ISIS-Communication-systems- Brochure-v2-compressed.pdf [11] Clyde Space, (2019) “CPUT S-Band Patch Antenna” [Brochure]. Retrieved from: https://www.clyde.space/products/13-cput-sband-patch-antenna [12] Agencia Espacial Civil Ecuatoriana (2017) “DSA 1U Deployable Solar Arrays” [Brochure]. Retrieved from:https://www.cubesatshop.com/wp-content/uploads/2016/07/EXA-DSA- Brochure-3C.pdf [13] GOM Space A/S, “NanoPower BP4”, DS 1013024 2.7 Datasheet, 2018. Retrieved from: https://gomspace.com/UserFiles/Subsystems/datasheet/gs-ds-nanopower-bp4-27.pdf [14] GOM Space A/S, (2018) “NanoCam C1U”, [Brochure]. Retrieved from: https://gomspace.com/UserFiles/Subsystems/flyer/NanoCamC1U_HIGH.pdf Project funded by the Interreg Sudoe Programme through the European Regional Development Fund (ERDF) 51 [15] Arianespace ArianeGroup. (2018). “Soyuz at the Guiana Space Centre”: User’s Manual. Kourou Cedex French Guiana: Arianespace. Retrieved from: http://www.arianespace.com/wp-content/uploads/2015/10/Soyuz-UsersManuel-issue2- Revision1-May18.pdf [16] Innovative Space Logistics (2019) “Launch services”, [Brochure]. Retrieved from: https://www.isispace.nl/wp-content/uploads/2019/01/ISL-Launch-services-brochure- compressed-web-version.pdf [17] Encinas. J.M. et al, “Xatcobeo: Small Mechanisms for CubeSat Satellites – Antenna and Solar Array Deployment”, Proceedings of the 40th Aerospace Mechanisms Symposium, NASA Kennedy Space Center, May 2010, NASA/CP-2010-216272 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20100021944.pdf [18] Nava, N. et al. “A novel hold-down and release mechanism for non-explosive Actuators based on SMA technology”, 16th European Space Mechanisms and Tribology Symposium, Bilbao, Spain, September 2015, ESA SP-737 http://esmats.eu/esmatspapers/pastpapers/pdfs/2015/nava.pdf [19] Tsay, M., et al., “Iodine-Fueled Mini RF Ion Thruster for CubeSat Applications,” 30th International Symposium on Space Technology and Science, Hyogo-Kobe, Japan, July 2015, ISTS-2015-b-273. [20] Tummala, A.R., Dutta, A. “An Overview of Cube-Satellite Propulsion Technologies and Trends” Aerospace 2017, 4(4), 58; doi: 10.3390/aerospace4040058 [21] Busek Co. Inc., “3cm RF Ion Thruster BIT-3,” 70010819 RevA Datasheet, 2014. Retrieved from: http://www.busek.com/index_htm_files/70010819%20RevA%20Data%20Sheet%20for%20 BIT-3%20Ion%20Thruster.pdf [22] Busek Co. Inc., “BET-1mN Busek Electrospray Thruster,” 70008500H Datasheet, 2016. [23] Enpulsion, “IFM-NANO Thruster”, ENP 2018-001 (Rev. E.2) Datasheet, 2018 [24] Action Systems, “TILE-5000 Modular Small Satellite Propulsion”, Datasheet. https://static1.squarespace.com/static/5446faa2e4b025843cfc6731/t/592f21edd1758edb9f7d bb8c/1496261102334/TILE-5000+Data+Sheet.pdf [25] ArianeGroup GmbH, (2018) “1N Hydrazine Thruster” [Brochure]. Retrieved from http://www.space-propulsion.com/spacecraft-propulsion/hydrazine-thrusters/1n-hydrazine- thruster.html [26] Tsay, M., et al., “Flight Development of Iodine BIT-3 RF Ion Propulsion System for SLS EM-1 CubeSats,” 30th AIAA/USU Conference on Small Satellites, North Logan, Utah, August 2016, SSC16-WK-39. [27] HTBST-ST200 Star Tracker. Retrieved from: https://hyperiontechnologies.nl/wp-content/uploads/2018/07/HTBST-ST200-V1.01_Flyer.pdf [28] New Space Sun Sensor- NCSS-SA05. Retrieved from: https://www.cubesatshop.com/product/digital-fine-sun-sensor/ Project funded by the Interreg Sudoe Programme through the European Regional Development Fund (ERDF) 52 APPENDICES APPENDIX A – REPORT CHANGE RECORD TABLE Edition/Revision Date Description of the change V0.0 22/04/19 Initial version of the document V1.0 03/05/19 Inclusion of Appendices Section Inclusion of Acronyms and References V1.1 08/05/19 Inclusion of risk analysis. V1.2 10/019 Final version of the document. Minor changes related to last document review TABLE 34 – PDR Document Change Record Project funded by the Interreg Sudoe Programme through the European Regional Development Fund (ERDF) 53 APPENDIX B – REQUIREMENTS SUBSYSTEM REQUIREMENTS CHANGE RECORD TABLE Requirements identification and development is the result of a continuous iterative process characterised by trade studies subsequently undertaken. As the requirements flow-down and the design matures, several updates on its initial definition are to be done. Table 35 gathers some of the major modification to which the different S/S requirements have been subjected to, as a brief exemplification of the updating practices that have been applied along the design phase. Edition/Revision Date Description of the change V0.0 25/02/19 Initial version of the document V1.0 03/03/19 CR(002-008) ER(002,003) TR(003) V1.1 10/03/19 CR(009-017) ER(004-) PR MeR(003-008), V2.0 20/03/19 CR(016), ER(004,009-012), PR(005) AR (003) MeR(003-005,007,009,011-014) TR(009-014) V2.1 22/04/19 CR(0016), ER(005,007,011), PR(001,004) AR (004,006,009,010) MeR(010,013) TR(009,014) V3.0 24/04/19 Final version of the document: TR(004-008,015) TABLE 35 - SS Requirement Change Log Record Project funded by the Interreg Sudoe Programme through the European Regional Development Fund (ERDF) 54 COMMUNICATIONS S/S Project funded by the Interreg Sudoe Programme through the European Regional Development Fund (ERDF) 55 Table 36 - Communications S/S Requirements ELECTRIC POWER S/S Project funded by the Interreg Sudoe Programme through the European Regional Development Fund (ERDF) 56 TABLE 37 – Electric Power S/S Requirements SPACE PROPULSION S/S TABLE 38 – Space Propulsion S/S Requirements Project funded by the Interreg Sudoe Programme through the European Regional Development Fund (ERDF) 57 ATTITUDE DETREMINATION AND CONTROL DETERMINATION S/S TABLE 39 – ADCS S/S Requirements Project funded by the Interreg Sudoe Programme through the European Regional Development Fund (ERDF) 58 STRUCTURE, MECHANICAL AND MECHANISIMS S/S TABLE 40 – Structure, mechanical and mechanisms S/S Requirements Project funded by the Interreg Sudoe Programme through the European Regional Development Fund (ERDF) 59 THERMAL CONTROL S/S REQUIREMENTS TABLE 41 – Thermal Control S/S Requirements Project funded by the Interreg Sudoe Programme through the European Regional Development Fund (ERDF) 60 APPENDIX C – MINUTES OF MEETINGS, MoM’s. TABLE 42 - MoM 1 Hereafter, the second taken decision related to the roles duties is to be further detailed: . Project Manager: Plan, schedule and coordinate the work to be done, monitoring the progress of team members. Track and refine the list of requirements. Take notes of the discussions/decisions/ next-actions and report them on the MoM’s. Ensure correctness and quality of the PDR. Contacting the NANOSTAR consortium representatives. (e.g. mails, Slack channels, etc.) . Telecommunications & Power Engineer: Identify Requirements affecting or motivated by this S/S. Sizing of communication link and definition of associated hardware. Selection of the GS to be used including antennas definition. Sizing of the power S/S of the space segment. Responsible of the power budget of the S/C. . Propulsion & ADCS Engineer: Identify Requirements affecting or motivated by this S/S. Estimation of the ΔV required for each phase of the mission. Select and size the propulsive system and associated hardware. Responsible for the ΔV budget of the space segment. Select and size the ACDS sensors and actuators. . S/C Mechanical design Engineer: Identify Requirements affecting or motivated by this S/S. Definition of the nominal configuration & attitude of the S/C in each mission phase. Definition of the structural S/S, materials, and mechanisms of the space segment. Computation of the center of mass location and inertia matrix of the S/C. Project funded by the Interreg Sudoe Programme through the European Regional Development Fund (ERDF) 61 Responsible for the mass and volume budget of the space segment. . Thermal control Engineer: Identify Requirements affecting or motivated by this S/S. Definition of a simplified S/C thermal model. Identify and limit the equipment temperatures. Check if they are fulfilled. Definition of passive/active thermal control techniques. Gather information on the space environment at each mission phase. Lastly, mention that the entire team has participate in the identification of the most important mission show-stopper and together propose possible mitigation strategies. Therefore, the responsibility is shared. TABLE 43 - MoM 2 Project funded by the Interreg Sudoe Programme through the European Regional Development Fund (ERDF) 62 TABLE 44 - MoM 3 Project funded by the Interreg Sudoe Programme through the European Regional Development Fund (ERDF) 63 TABLE 45 - MoM 4 TABLE 46 - MoM 5 Project funded by the Interreg Sudoe Programme through the European Regional Development Fund (ERDF) 64 TABLE 47 - MoM 6 Project funded by the Interreg Sudoe Programme through the European Regional Development Fund (ERDF) 65 APPENDIX D – TRADE-OFF STUDIES As a support material for the trade-off analysis that have been explained along section 3, this appendix gathers the criteria used for the definition of the weighting factors as well as the different parameters taken into account for each subsystem. Moreover, for those trade-off analysis in which just the qualitative results were provided, here the assigned factors of each of the compared design choices are presented justifying in a further extent the selected designs. 1) Space Propulsion S/S: TRANSFER POWER DELTA-V NUMBER OF DRY MASS GRADING DURATION (W) (KM/S) THRUSTERS (KG) (MONTHS) 5 <20 <2.5 <12 1 >5 4 20 – 25 2.5 - 2.75 12 – 24 2 4.5 – 5 3 25 – 30 2.75 - 3 24 – 36 3 4 – 4.5 2 30 – 35 3 - 3.25 36 – 48 4-5 3.5 – 4 1 35- 40 3.25 - 3.5 48 – 54 6-8 3 - 3.5 0 >40 >3.5 >54 >8 <3 WEIGHTING 4 2 3 4 5 FACTOR (1-5) TABLE 48 - Evaluation Criteria for Propulsive Trade-Off Analysis 푆푐표푟푒 = ∑ 푊푒𝑖푔ℎ푡𝑖푛푔 푓푎푐푡표푟 · 푔푟푎푑𝑖푛푔 퐶푎푡푒푔표푟𝑖푒푠 NOTE: if a 0 is obtained in any of the categories, the total score shall be 0 since this would imply that the solution does not meet the requirement. 2) Telecommunications S/S: EVALUATION CRITERIA UHF BAND S-BAND (Weight and Parameter) Value Assigned (1-5) Value Assigned (1-5) (Weight 1-8) Mass (2) 2 2 Cost (2) 2 3 Gain (5) 1 4 Power Consumption (8) 1 5 Installability (4) 2 3 Thermal Performance (5) 3 2 Total Mark 44 92 (Weight X Value) TABLE 49 - Evaluation Criteria and Results for Telecommunications Trade-Off Analysis Project funded by the Interreg Sudoe Programme through the European Regional Development Fund (ERDF) 66 3) Attitude Control and Determination S/S: Sun Star Horizon Weight Parameters IMU GNSS Magnetometers Sensors Sensors Sensors 5 Cost 0 - 5 1 - - 10 Quality for the mission 10 0 7 10 0 0 5 Mass 0 - 10 5 - - 4 Power Consumption 0 - 5 0 - - 10 Accuracy 10 - 5 10 - - FINAL RATE 200 215 230 TABLE 50 - Weighted Parameters for ADCS sensors Trade-Off Analysis Cold Gas Liquid Rockets BUSEK Main Thruster to Weight Parameters MEMS Hydrazine Thruster BIT-1 Desaturate 5 Cost 5 5 5 10 10 Quality for the mission 8 8 8 8 10 Mass 4 6 8 10 10 Accuracy 7 7 7 7 FINAL RATE 215 235 255 300 TABLE 51 - Weighted Parameters for ADCS Desaturation system Trade-Off Analysis Weight Parameters RWP050 RWP100 RWP500 RW1 5 Cost 10 8 6 4 10 Quality for the mission 7 7 7 7 10 Mass 10 8 6 4 10 Power Consumption 10 8 6 4 7 Accuracy 10 8 6 4 FINAL RATE 390 326 262 198 TABLE 52 - Weighted Parameters for ADCS RW Trade-Off Analysis 4) Mechanical, Structure and Mechanisms S/S: WEIGHT PARAMETERS CONFIGURATION 1 CONFIGURATION 2 5 Mass 3 5 4,5 Inertia 1 4 4 Volume 3 4 3,5 Distribution 4 5 4 Mechanisms Accuracy 4 4 FINAL GRADE max = 105 61,5 92,5 TABLE 53 – Weighted Parameters for Mechanical Trade-Off Analysis Project funded by the Interreg Sudoe Programme through the European Regional Development Fund (ERDF) 67 APPENDIX E – RISK ANALYSIS/ MITIGATION TABLE Following what has been explained at Section 4 of the present document, hereafter the found risks are presented. They are related to the mission accomplishment itself and the performance of the subsystems conforming it. Detailed information is provided at the following table where the cause of each risk, its description, the planned response strategy to be followed and the way those actions are to be executed, are all summarised in accordance with the identified level of severity. Notice that the owner of the risk has also been defined in order to define the responsible of studying its effects and mitigation process. RESPONSE TITLE CAUSE DESCRIPTION OWNER ACTION SEV. STRATEGY Increment of total mass or COG deviation can Smaller COG DAMAGE Coupling of modify the inertia, Mechanical deviation margin DURING natural Mitigate 4 leading to a lower Subsystem must be set in the LAUNCH frequencies natural frequency of the PDR system Solar array Bolt supporting Increase the batteries FAILURE OF THE deployment Mechanical mechanical limits is lost Transfer as secondary power 4 POWER SUPPLY mechanism does Subsystem by impact breakage source not work properly The adjacent antenna FAILURE OF Bending of the support structure Mechanical Avoid Tolerances 3 SENSORS structure deformation hinders Subsystem the sensor operation DAMAGE Structure ADCS accuracy must Perturbations or solar Mechanical DURING frequencies Transfer avoid the appearance 3 arrays induced torques Subsystem OPERATION resonance of parasitic torques Certificate the FAILURE OF THE Deployable mechanism Mechanical mechanism by means Mechanical failure Avoid 5 PAYLOAD does not work properly Subsystem of severe test campaign REACTION Fatigue Rupture Bearing failure causes WHEEL (exceedance in Start using the three reaction wheel to AOCS Mitigate 1 BEARING bearing lifetime remaining wheels become non operative FAILURE cycles) or friction Loss of absorbed water LACK OF Outgassing of Start using the three due to the very low AOCS Mitigate 1 LUBRICATION lubricants remaining wheels pressures UNDER- Use high margins DIMENSION OF Incorrect safety Time of desaturation reducing the WHEEL margins and higher than the time AOCS Avoid 1 possibility of MOMENTUM estimations between desaturations occurrence STORAGE UNDERDIMENSI Use high margins Incorrect safety Torque manoeuver ON THE PEAK reducing the margins and cannot be handled by AOCS Avoid 1 MOMENTUM OF possibility of estimations the system THE WHEEL occurrence SENSOR High-Energy particle Start using back up FAILURE DUE TO Pass through Van radiation can penetrate AOCS Mitigate sensor to control 1 HIGH CHARGED Allen Belts and charge dielectrics AOCS PARTICLES Incorrect input RCS THRUSTER power, Loss of a RCS thruster Ceasing system SYSTEM AOCS Avoid 0 mechanism for desaturation development FAILURE failure etc. MAIN THRUSTER Evaporating Propellant does not AOCS Accept Risk acceptance 5 FAILURE system failure. reach the thruster. A larger Iodine tank Perform test for may affect the LOWER different tank Different Iodine sublimation Propulsion PROPELLANT Avoid dimensions position 4 tank size process/system, leading engineer EFFICIENCY of the sublimation to a lower propellant thermistor(s) efficiency Project funded by the Interreg Sudoe Programme through the European Regional Development Fund (ERDF) 68 Perform the targeting Trajectory Trajectory requirement analysis to calculate DELTA-V Correction maneuvers are several Propulsion the specific TCM to be MARGIN IS Avoid 3 Maneuvers require more/longer than engineer performed in order to INSUFFICIENT more propellant initially assumed reduce the uncertainty Ensure thickness of TEMPERATURE Heat radiated by the propellant tank is OF THE sun or other S/C Propellant tank is Propulsion enough to provide the PROPELLANT components is exposed to a too large Mitigate 3 engineer required insulation TANK IS OUT OF insufficient/too or small temperature (thermoplastic LIMITS high material) Cathode and thruster NON- may not produce the Tests to certify the QUASINEUTRAL Cathode particles same amount of Propulsion Mitigate thruster shall ensure 2 PLUME OF THE not well regulated particles to ensure engineer this risk is minimum THRUSTER quasi-neutrality of the plume Not using the subsystems at the All systems working at same time, trying to full power at the same CRITICAL space out their usage Contingency time, thus an PE Acceptance 3 SCENARIO in order to diminish unpredicted increment the power of power takes place consumption at each instant Consider that any type of failure of the batteries are fault of The battery is not the supplier and shall FAILURE OF A Overheating, loading enough power, PE Transfer have been considered 2 BATTERY cabling system critical during eclipses. for its manufacturing. Thus, the engineering team is not responsible. The solar array is stuck Through tests on in an inconvenient SOLAR PANEL ground and correct Failure of position due to DEPLOYMENT PE Mitigation installation of the 3 mechanism mechanism failure ERROR solar arrays, checked during the deployment with tests. process. Solar arrays are not generating enough Locating the solar power at EOL due to a EXAGERATED panels far from the higher internal DEGRADATION Radiation and thruster and selecting degradation than the PE Avoidance 2 OF SOLAR temperature cells with low expected theoretically. ARRAYS degradation The radiation can be probability. due to thruster emission of particles. The batteries are not Using buses that LOW VOLTAGE loaded due to the low allow converting the OF SOLAR voltage of the current voltage and assure Inconvenient ARRAY coming from the solar PE Avoidance that the input current 1 design choice GENERATED array. Thus, the batteries to the battery has an POWER are not able to load acceptable current power. value. The cables are not Tests performed on connecting correctly CABLING ground to assure the Incorrect design the equipment of the PE Mitigation 1 FAILURE correct connectivity subsystem and between equipment. between subsystems. Simulations on ground that allow The power supplied by assuring the power UNDERSIZING the batteries is not level of the batteries OF THE Incorrect design PE Mitigation 2 enough to feed the and the required BATTERIES subsystems. power that shall be provided during eclipses. Project funded by the Interreg Sudoe Programme through the European Regional Development Fund (ERDF) 69 The supplier of the Several subsystems are voltage conversor not provided with the INCORRECT buses is responsible power that they require POWER Incorrect design PE Transfer for its correct 2 due to an inefficient DISTRIBUTION functioning assured distribution provided by once the equipment the voltage buses. is bought. The radiation and debris of the free-space A margin on the link ANTENNA Environment damage the surface of budget was set to SURFACE TE Acceptance 2 contingencies the antenna, increasing account for this type DEGRADATION the power for the of contingencies. transmission A possible change in The downlink is the trajectory after the seized so the flyby can impose a CHANGE IN THE Mission transmission can be smaller transmission TE Acceptance 3 ORBIT contingencies performed under two time than the required, days in case of this thus the data link will type of contingencies. not be completed. LOW VISIBILITY Any type of change in The mission analysis OF THE Mission the orbit from the engineer shall assure TE Avoidance 3 GROUND contingency expected could lead to a very low probability STATION a poor transmission line. of trajectory change. Possible failure of the The correct FAILURE OF THE data handling and functioning of the DATA Wiring, power communications device TE Transfer device shall be 1 HANDLING supply due to bad wiring or assured by the COMPUTER lack of enough power supplier. supply. Failure of downlink transmission after flyby The correct ERROR IN THE due to an error in the functioning of the Malfunction of the UPLINK/DOWNLI transceiver that does TE Transfer device shall be 1 transceiver NK SWITCH not permit the switch to assured by the downlink supplier. communications. TABLE 54 - Risk Analysis / Mitigation Project funded by the Interreg Sudoe Programme through the European Regional Development Fund (ERDF) 70 APPENDIX F – LAUNCH PHASE DESCRIPTION The Soyuz three main phases are briefly described 1) Ascent of the Soyuz three-stage: The Fregat upper stage is a restartable upper stage (up to 7 times), offering a great flexibility to servicing a wide range of orbits and allowing delivering the payload to different orbits in case of shared launch, which is considered to be our mission case, as the CubeSat is to be very small compared with the possible principal payload of the launcher. At the end of the three-stage Soyuz phase, the upper composite (Fregat with payload) is separated on a sub-orbital path. FIGURE 28 - Schematic Soyuz Launch [15]. 2) Fregat upper stage flight profile for payload delivery to final orbit: In the present case, for elliptic equatorial orbit, including GTO, a single Fregat boost injects the upper composite into the targeted orbit (direct ascent profile). As it can be seen in the picture, the mission duration is around 30 mins (from lift off to separation). FIGURE 29 - Timeline of Soyuz Launch Sequence [15] Project funded by the Interreg Sudoe Programme through the European Regional Development Fund (ERDF) 71 3) Fregat deorbitation or orbit disposal maneuvers. FIGURE 30: SCHEMATIC GTO- REACHING.[15] Launch window Using ISL launch services, a possible launch window is Q2 2020 from the advantageous location of the Guiana Space Centre. The site low latitude minimizes the satellite on-board propellant needed to reach the equatorial plane, providing additional lifetime so provided that the detailed mission design for the nanosatellite ends up with that needed inclination, this launch window shall be considered. Orientation and separation conditions After injection into the orbit, the Fregat Attitude Control System (ACS) is able to orient the upper composite to any desired attitude and to perform separation in 3-axis stabilization mode. In case the maximum spacecraft static unbalance remains below 15 mm, the typical spacecraft 3σ pointing accuracies after S/C separation, for a three-axis stabilised mode are: - Geometrical axis depointing ≤ 4 deg - Angular tip-off rates along longitudinal axis ≤ 1 deg/s - Angular tip-off rates along transversal axes ≤ 1.5 deg/s Project funded by the Interreg Sudoe Programme through the European Regional Development Fund (ERDF) 72 FIGURE 31- Typical Sequence of events during Payload deployment Stabilization Regarding the separation linear velocities and collision risk avoidance, it is relevant to mention that the payload adapter’s separation systems are designed to deliver a minimum relative velocity between spacecraft and upper stage of 0.5 m/s. Launch global mechanical environment - Quasi static Load: The highest longitudinal acceleration occurs just before the first-stage cut- off and does not exceed 4.3 g. The highest lateral static acceleration may be up to 0.4 g at maximum dynamic pressure and takes into account the effect of wind and gust encountered in this phase. - Line peak loads: Such local over line loads are specific of the adapter design. For off-the-shelf adapters, a value of 15% over the average line loads seen by the spacecraft is to be taken into account. - Sine vibration loads: in the frequency of 60-100Hz, both lateral and longitudinal maximum sine loads are 0,3g in the s/c-adapter interface. - Root mean square vibration maximum level (GRMS) occurs during the first stage flight and its value is 4.94 grms. - During flight, acoustic pressure fluctuations under the fairing are generated by engine plume impingement on the pad during lift-off and by unsteady aerodynamic phenomena during atmospheric flight. So, it is recommended that Spacecraft envelope volume ≤ 40 m3 Cross- section ≤ 6.4, when the filling factor is ≤50%. The shock due to the separation at the interface is 20g at 100Hz, and 1000g when it reaches 1000Hz. Project funded by the Interreg Sudoe Programme through the European Regional Development Fund (ERDF) 73 APPENDIX G – DESIGN AND STRUCTIRE DETAILS FIGURE 32 - Propulsion Subsystem FIGURE 33 - Power Subsystem Project funded by the Interreg Sudoe Programme through the European Regional Development Fund (ERDF) 74 FIGURE 34 - ADCS Subsystem FIGURE 35 - OBC Project funded by the Interreg Sudoe Programme through the European Regional Development Fund (ERDF) 75 FIGURE 36 - Communications Subsystem FIGURE 37 - Structure Subsystem Project funded by the Interreg Sudoe Programme through the European Regional Development Fund (ERDF) 76 Mechanisms for the Launcher-satellite interface The Soyuz offers a range of standard off-the-shelf adapters and their associated equipment, compatible with most of the spacecraft platforms. The adapter and carrying structure within the launcher interface are depicted below. PAS 937 ADAPTER: A clamp-band with a ASAP-S CARRYING STRUCTURE: the low shock separation system (CBOD) is in separation system allows to jettison the charge of this operation.[15] upper part ring by low shock separation (EADS CASA 8 springs).[15] ISIS 6 UNIT DUOPACK: Ultra-light burnwire based Hold Down and Release Mechanism. The proposed mechanism a Pin Puller and a Hold-Down and Release Actuator (HDRA) called REACT (REsettable non-explosive ACTuator).[7] Project funded by the Interreg Sudoe Programme through the European Regional Development Fund (ERDF) 77 ConfigurationAPPENDIX : H – CONFIGURATIONScfg.LAUNCH Configuration MASS LAUNCH BUDGET ▼ NANOSAT Target wet mass [Kg] : + Unit Without margin [Kg] Margin [%] Margin [Kg] Including margin [Kg] % of total Subsystem - Name Quantity Mass [Kg] Margin [%] ▼ Subsystem PROP 1,70 0,00% 0,00 1,70 27,11% ▼ Ion Thruster system 1 1,66 0,00% 1,66 0,00% 0,00 1,66 | BIT3-with 2ax gimbals 1 1,44 0,00% 1,44 0,00% 0,00 1,44 | Prop control valve 1 0,11 0,00% 0,11 0,00% 0,00 0,11 | Tank pressure transducer 1 0,02 0,00% 0,02 0,00% 0,00 0,02 | Iodine tank 1 0,09 0,00% 0,09 0,00% 0,00 0,09 Power generator board+PPU 1 0,05 0,00% 0,05 0,00% 0,00 0,05 ▼ Subsystem PAY 1,00 5,48% 0,05 1,05 16,82% Gomspace - NanoCam 35mm 1 0,97 5,00% 0,97 5,00% 0,05 1,02 Cover 1 0,03 20,00% 0,03 20,00% 0,01 0,04 Insulator pad 1 0,00 20,00% 0,00 20,00% 0,00 0,00 ▼ Subsystem STR 0,54 20,00% 0,11 0,65 10,35% ▼ Tutorial Custom 3U 1 0,54 20,00% 0,54 20,00% 0,11 0,65 ▼ Subsystem PWR 0,99 5,00% 0,05 1,04 16,58% Gomspace - NanoPower BP4 F Option1 0,27 5,00% 0,27 5,00% 0,01 0,28 ▼ Solar Array 2 0,36 5,00% 0,72 5,00% 0,04 0,76 | ▼ Assembly 13 1 0,36 5,00% 0,36 5,00% 0,02 0,38 | | ▼ 2U Generic Deployable Solar 2Panel 0,18 5,00% 0,36 5,00% 0,02 0,38 | | | 2U Generic Solar Panel 2 0,09 5,00% 0,18 5,00% 0,01 0,19 | Equipment 15 1 0,00 5,00% 0,00 5,00% 0,00 0,00 ▼ Subsystem ADCS 1,17 15,00% 0,18 1,35 21,45% Hyperion Tech - HT RW210.15 4 0,24 15,00% 0,96 15,00% 0,14 1,10 Auriga Star Tracker Sodern 1 0,21 15,00% 0,21 15,00% 0,03 0,24 ▼ Subsystem OBC 0,27 15,00% 0,04 0,31 4,95% ISIS Computer with Daughter Board 1 0,27 15,00% 0,27 15,00% 0,04 0,31 ▼ Subsystem COMM 0,13 0,00% 0,00 0,13 1,99% ISIS UHF/VHF Transceiver 1 0,08 0,00% 0,08 0,00% 0,00 0,08 ClydeSpace - CPUT S-Band Patch Antenna1 0,05 0,00% 0,05 0,00% 0,00 0,05 ▼ Subsystem THERMAL 0,04 20,00% 0,01 0,05 0,75% Radiator + Al-cover 1 0,04 20,00% 0,04 20,00% 0,01 0,05 Total dry mass without system margin 5,84 7,48% 0,44 6,27 System margin 0,00% 0,00 6,27 Propellant mass 2,73 15,00% 0,41 3,14 Total wet mass including all margins 9,41 SuperCollapser:End System Without margin [Kg] Margin [%] Margin [Kg] Including margin [Kg] Total dry mass without system margins 5,84 0,44 6,27 Total dry mass including system margins 6,27 Total propellant mass 2,73 15,00% 0,41 3,14 Total wet mass including all margins 9,41 TABLE 55 - Launch Configuration Mass Budget As it can be seen, the subsystem that contributes the most is the propulsion subsystem, as it could be expected from the Iodine tank mass, followed by the ADCS subsystem, composed by heavy components as the 4 RW and the star tracker. This is the principal reason for the cg. slight deviation from the cross-sectional area in the x axis. It is as well noticeable that the lighter subsystem is the Communications one, which mainly gathers low mass components. Project funded by the Interreg Sudoe Programme through the European Regional Development Fund (ERDF) 78 TABLE 56 - Transfer Mode Mass Budget Project funded by the Interreg Sudoe Programme through the European Regional Development Fund (ERDF) 79 TABLE 57- Science Mode Mass Budget Project funded by the Interreg Sudoe Programme through the European Regional Development Fund (ERDF) 80 APPENDIX I – ADDITIONAL TELECOMMUNICATIONS S/S INFORMATION KOUROU STATION 3 FIGURE 38- Kourou Station 3 https://m.esa.int/Our_Activities/Operations/Estrack/Kourou_station Project funded by the Interreg Sudoe Programme through the European Regional Development Fund (ERDF) 81 FIGURE 39 – Selected Antenna Gain Evolution Project funded by the Interreg Sudoe Programme through the European Regional Development Fund (ERDF) 82 APPENDIX J – ADDITIONAL INFORMATION OF THE THERMAL S/S Orbital scenarios FIGURE 40 - Orbital Results for the Eclipse (At Perigee) Scenario FIGURE 41 - Orbital Results for the Sun Light Scenario Project funded by the Interreg Sudoe Programme through the European Regional Development Fund (ERDF) 83 FIGURE 42 - Orbital Results for the Hypothetic Scenario With Eclipse At Apogee Components temperature evolution FIGURE 43 - Nodes Temperature Evolution for the Eclipse (At Perigee) Scenar io Project funded by the Interreg Sudoe Programme through the European Regional Development Fund (ERDF) 84 FIGURE 44 - Nodes Temperature Evolution for the Eclipse at Apogee Scenario FIGURE 45 - Nodes Temperature Evolution for the Sun Light Scenario Project funded by the Interreg Sudoe Programme through the European Regional Development Fund (ERDF) 85