Preliminary Design of a Global and Continuous Coverage Communication Services Constellation

Transcription

1 CRANSEDS - CRANFIELD UNIVERSITY UKSEDS-SSPI 2016 Satellite Design Competition Preliminary Design of a Global and Continuous Coverage Communication Services Constellation June 2017

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3 ABSTRACT The following document contains a precise description of the developed preliminary design of a constellation of small satellites, and their composing satellites, which goal is to deliver global and continuous coverage. This document is submitted to UKSEDS and was developed by team, representing Cranfield University. The document was generated in response of and to participate in SSPI Satellite Competition. Keywords:, Constellation, Polar Orbits, Global Coverage, Continuous Coverage, Satellite Communications, Ka Band, GEO Interferences, Phased Array Antenna i

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5 TABLE OF CONTENTS ABSTRACT... i LIST OF FIGURES... vi LIST OF TABLES... viii LIST OF ABBREVIATIONS... x 1 Introduction Previous Missions Baselines Iridium Globalstar O3b OneWeb Mission Statement Mission Baseline Mission Requirements Budgets Initial Estimations Final Budgets and Comparison Cost Analysis Estimating Mission Lifetime Final Cost Estimation Satellite Constellation Constellation Requirements and Constraints Initial Concepts & Considerations Design of Final Constellation Requirements and Constraints on Constellation, Including Management of Interference Interference Mitigation Communication Subsystem Introduction Communications Links Architecture Inter-Satellite Links Gateway Stations Communication Payload Link Budget Frequency Data Rate Orbit Regulation Aspect and Interferences Mitigation Landing Rights and Spectrum Management Propulsion Subsystem iii

6 6.1 Engine Selection Options Airbus 200 N Bipropellant Thruster Fuel Storage Propulsion Subsystem Summary Launch and Orbit Launcher Selection Selection Method Atlas V 500 Series Parameters Launch Procedures Insertion into final Orbit Computation Method Station Keeping and Space Debris Delta-V and Propellant Budgets Summary End of Mission Considerations Disposal Options and Requirements Constellation Disposal Disposal Method Computation Attitude Determination and Control System ADCS System Overview ADCS Modes Detumbling and Data Acquisition Mode Normal Mode Orbit Correction Mode Safe Mode Design Considerations Hardware Selection Reaction Wheels Thrusters Star Trackers Sun Sensors Gyroscope Electrical Power Subsystem Power Requirements Power Budget Power Generation Primary Power Secondary Power Power Distribution, Management and Control EPS Mass Budget iv

7 11 On-Board Data Handling Subsystem Requirements OBDH Design Architecture & Hardware Memory Protection and Fault Tolerance Subsystem Interfacing Structure and Configuration Introduction Subsystem Requirements SSTL-150 Satellite Platform Structure Introduction Configuration Material Configuration Introduction Mechanisms Introduction Solar Array Deployment Mechanism Thermal Control Subsystem Mission Drivers for Thermal Design Overall Mission Requirements Thermal Requirements Thermal Modelling Conclusion Thermal Design Further Development REFERENCES APPENDICES Appendix A Atlas V 500 Series Launch System Appendix B Solar Cell Datasheet Appendix C Battery Datasheet v

8 LIST OF FIGURES Figure 1-1 Iridium satellite [1]... 2 Figure 1-2 Globalstar coverage [2]... 4 Figure 1-3 Globalstar satellite [2]... 5 Figure 1-4 O3b satellite [3]... 6 Figure 1-5 OneWeb satellite [4]... 8 Figure 1-6 Interferences with GEO satellites... 9 Figure 1-7 Progressive pitching used for avoiding interferences with GEO satellites... 9 Figure 1-8 satellite design Figure 1-9 Final mass budget percentages Figure 1-10 System mass comparison between initial estimation and preliminary design mass Figure 2-1 Average NASA small spacecraft mission Figure 3-1 Final constellation Figure 3-2 Orbit plane intersections Figure 3-3 Polar constellation Figure 3-4 Polar targeting Figure Walker delta with square beams Figure Walker delta with circular beams Figure 3-7 Minimum configuration with 36 RAAN spacing Figure 3-8 Minimum Configuration with 18 RAAN spacing Figure 4-1 Satellite communication link architecture [10] Figure 4-2 Classical satellite communication system [11] Figure 4-3 Intersatellite link between the planes [12] Figure 4-4 Transparent and regenerative repeaters [10] Figure 4-5 Link budget picturing [13] Figure 4-6 Spacecraft line sight geometry [14] Figure 6-1 Airbus 200 N Bipropellant thruster sketch vi

9 Figure 6-2 Propulsion subsystem design Figure 7-1 Hohmann transfer sampling points Figure 7-2 Spatial density > 10 cm (extracted from Operational Collision Avoidance by ESA Space Debris Office presentation given by Klaus Merz on 03/11/2016) Figure 8-1 IDAC protected regions (IADC-02-01, 2007) Figure 8-2 Orbital decay Figure 10-1 Li-Ion batteries [19] Figure 10-2 SST Power conditioning and distribution units [20] Figure 11-1 Example of OBDH interface [21] Figure 12-1 SSTL Figure 12-2 Basic CAD model of the structure Figure D views with dimensions of the structure Figure 12-4 CAD model of the satellite Figure 12-5 CAD model with systems breakdown Figure 12-6 View of the CAD model of the satellite Figure 12-7 Double fold and roll-up solar array. Image from the University of Cambridge Figure 13-1 Fluxes impacting LEO satellite Figure 13-2 Physical characteristics Figure 13-3 Heater system Figure 13-4 Louver system Figure 13-5 Multi-layer insulation vii

10 LIST OF TABLES Table 1-1 Initial mass budget Table 1-2 Initial power budget Table 1-3 Final mass budget Table 2-1 Estimated cost of the mission Table 3-1: Satellites and planes required at different altitudes Table 4-1 Link budget calculations Table 6-1 Thrusters considered Table 6-2 MRE-1.5 & 200 N efficiency comparison Table 6-3 Airbus 200 N Bipropellant thruster technical specifications Table 6-4 MOOG-ISP fuel tanks considered Table 7-1 Propellant mass for orbital boost Table 7-2 Atlas V 500 series launch capabilities into 1900 x 1900 polar orbit.. 56 Table 7-3 Constellation deployment procedure Table 7-4 Delta-V and propellant mass budgets Table 8-1 Constellation disposal trade-off Table 9-1 ADCS system requirements Table 9-2 ADCS hardware properties part Table 9-3 ADCS hardware properties part Table 10-1 Power budget Table 12-1 Main characteristics of the SSTL Table 12-2 External dimensions of the structure Table 12-3 Structure breakdown Table 12-4 Internal volume calculation Table 12-5 Available internal dimensions Table 12-6 Aluminium-skinned honeycomb main properties Table 12-7 Total mass and surface implemented in CATIA V Table 12-8 Inertia matrix calculated by CATIA V viii

11 Table 13-1 Functional requirements of the thermal control subsystem Table 13-2 Temperature requirements for each subsystem Table 13-3 Summary of critical phases, distances and power dissipated Table 13-4 Material characteristics Table 13-5 Physical characteristics Table 13-6 Estimated average temperature Table 13-7 Satellite temperatures per mission phase Table 13-8 Thermal control subsystem power and mass budgets ix

12 LIST OF ABBREVIATIONS AOCS CDMA Attitude and Orbit Control Systems Code Division Multiple Access Cranfield Students for the Exploration and Development of Space CU EPS ESA FAT FSS GEO GTO IDAC ITU KE LEO MEO MLI MMH MSS NASA OBC OBDH PCDU PLF RAAN SBD SSPI TT&C ULA Cranfield University Electrical Power System European Space Agency Frequency Allocation Table Fixed Satellite Service Geostationary Orbit Geosynchronous Transfer Orbit Inter-Agency Space Debris Coordination Committee International Telecommunication Union Kinetic Energy Low Earth Orbit Medium Earth Orbit Multi-Layer Insulation MonoMethyl Hydrazine Mobile Satellite Service National Astronautics and Space Administration On-board Computer On-Board Data Handling Power Conditioning and Distribution Unit PayLoad Fairing Right Ascension of the Ascending Node Short Burst Data Society of Satellite Professionals International Telemetry Tracking & Commend United Launch Alliance x

13 Introduction 1 Introduction This section introduces the performed design developed by the team, from Cranfield University, participating in the UKSEDS-SSPI 2016 Satellite Design Competition. In this first subsection section is explained current and past projects based on constellations of communication satellites, the mission baseline, requirements of the mission, and budgets discussion. Following subsections will explain the mission statement, baseline and requirements, concluding with the initial and final budgets. 1.1 Previous Missions Baselines Here is presented a group of four space missions performing global communication activities with constellations. They are presented in chronological order Iridium Main Mission Parameters Spacecraft mass 690 kg Spacecraft power 400 W Spacecraft lifetime 5 years Global coverage Yes Continuous coverage No Number of satellites 66 (+6) Orbit altitude 780 km Iridium mission is operated by a USA company. This is the greatest current constellation and continue under evolution. For example, their new generation constellation Iridium NEXT is expected to be delivered by 2018 [1]. Its constellation is formed by 66 operational satellites, plus 6 more on-orbit spares, distributed in 6 polar orbit planes. The designed orbit altitude is 780 km, with an inclination angle of 86.4 degrees and 8.2 degrees of minimum elevation angle [1]. This constellation provides global but not continuous coverage. 1

14 Introduction Iridium service is based on user terminals communication systems. Satellites communicate with them using an L band at MHz, enabling a telephony and modem data rate of 2.4 kbps. Each satellite has 4 inter-satellite links operating in Ka band at GHz and communicate with Ground stations using Ka band links at GHz (uplink) and GHz (downlink). These ground stations are called Iridium Gateways and connect the Iridium network with the ground stablished telephone network [1]. Figure 1-1 Iridium satellite [1] User terminals are Iridium's short burst data (SBD) transceivers that transmit data at L band with satellites and communicate with a range of Iridium mobile phones [1]. They provide voice and data service coverage. SBD transceivers are 400 g weight with an antenna length of 15 cm and mean and peak transmission power values of 0.6 and 7 W respectively. The following information of each spacecraft subsystem is obtained from the main webpage of the mission [1]: The communication subsystem uses four gimbaled nadir-pointed Ka band antennas to transmit and receive to and from gateways. Two gimbaled Ka 2

15 Introduction band antennas are used for East-West communications, in other words, to communicate between orbital planes, and another two fixed Ka band antennas are used to communicate across the same orbit plane (North- South). Finally, 3 deployable phased antennas are used for the L band user link. The power subsystems use GaAs solar arrays of 3.9 m 2 and NiH2 batteries. Iridium s satellite is 3 axes stabilised using momentum wheels, magnetorquers, and thrusters. To close the control loop, they use line horizon sensors, three axis gyros and magnetometers. Orbit manoeuvres are performed with a redundant system of monopropellant hydrazine propulsion systems: a single electro-thermal hydrazine thruster and seven hydrazine reaction engines assemblies. The nominal fuel load of the system is kg. Thermal control system is passive with electronically controlled radiators. They use blankets and radiators. Finally, the structure system is based in a graphite epoxy triangular monocoque and truss structure. Solar array deployment systems are nonexplosive, however, launch separation mechanism is pyro actuated Globalstar Main Mission Parameters Spacecraft mass 700 kg Spacecraft power 1100 W Spacecraft lifetime 15 years Global coverage No polar coverage Continuous coverage No Number of satellites 32 Orbit altitude 1,400 km This project is compound of 24 second generation satellites and 8 first generation satellites. These satellites are distributed along 8 orbital planes in 52 degrees inclined orbits [2]. Thus, this project does not provide global coverage since poles are not covered. Operational orbit altitude is 1400 km. 3

16 Introduction Coverage is focused on North America, Europe, Japan and Australia. It can be seen in Figure 1-2. Globalstar service is provided acting as a backup coverage to users that do not have ground stations access. A Globalstar s satellite receive the user signal, which can be voice or data, and is sent to the nearest ground station by the same satellite. Then, ground network sends the information to the final user. This Globalstar service can also be used if a ground station signal is lost to jump over this gap in the network [2]. Figure 1-2 Globalstar coverage [2] Communication is performed in S band achieving 1 Mbps downlink and 256 kbps uplink for each user in second generation satellites. In total, up to 1248 different users can be covered by each satellite. Each satellite deliver 16 different channels divided in 78 divisions using CDMA [2]. The following information of each spacecraft subsystem is obtained from the main webpage of the mission [2]: The communication subsystem uses S bands for service or user links, achieving a total data rate of 1.2 Gbps. C band is used for TT&C. 4

17 Introduction The power subsystems use tracking solar panels and batteries for eclipse and peak periods. AOCS implemented is based on 3 axis stabilisation. Utilised actuators are momentum wheels, magnetometers and GPS. To close the control loop is utilised sun and earth sensors. Orbit manoeuvres are performed with a monopropellant hydrazine based propulsion system. The nominal fuel load of the system is 76.5 kg. Figure 1-3 Globalstar satellite [2] O3b Main Mission Parameters Spacecraft mass 700 kg (450 kg dry) Spacecraft power 1000 W (EOL) Spacecraft lifetime 10 years Global coverage No Continuous coverage No Number of satellites 12 Orbit altitude 8,060 km O3b mission goal is to deliver satellite internet services and mobile backhaul services to emerging markets. 5

18 Introduction 12 satellites form a constellation fully scalable to meet market demands. They are found in equatorial orbits with approximately zero degrees of inclination providing a standard coverage around +/- 45 degrees latitude. The orbit height is 8,062 km, MEO, and the mission lifetime is 10 years. Satellites are using Ka-Band payload designed to enable the high-speed flow of data between locations on the ground. Twelve fully steerable antennas ensure an optimised connection to the area where data is needed. Two of them are used to connect with gateways and the other ten to connect with users. The whole constellation delivers 70 remote beams, each of which has a coverage of 700 km. Each beam is capable of delivering up to 1.6Gbps of data. Figure 1-4 O3b satellite [3] The following information of each spacecraft subsystem is obtained from the main webpage of the mission [3]: The communication subsystem uses 12 Ka bands steerable antennas. The power subsystems use two deployable three-segmented Gallium Arsenide solar arrays and a Lithium ion battery for storage. Solar arrays generate 1,700 W BOL and 1,000 W EOL. AOCS implemented is based on 3 axis stabilisation, provided by a combination reaction wheels and magnetorquers. Attitude determination 6

19 Introduction is provided by earth and sun sensors in conjunction with an inertial measurement unit. Propulsion system applies a hydrazine monopropellant system, compound of 8 thrusters, with 141 kg of fuel. OBDH implement a LEON3 microprocessor and a MIL-STD-1553B Data Bus connecting all systems to the computer. The structure of the satellite is a trapezoidal in shape, consisting of rigid aluminium honeycomb panels OneWeb Main Mission Parameters Spacecraft mass 150 kg Spacecraft power Unknown Spacecraft lifetime Unknown Global coverage Yes Continuous coverage Yes Number of satellites 648 to 882 Orbit altitude 1,200 km OneWeb is a mission under development that is being currently designed and that will provide global and continues coverage to deliver internet access in every part around the globe. A few technical information has been released from the OneWeb mission and satellites. It is known that the constellation is wanted to be fully operable by 2027 with an estimated investment of more than 3 billion dollars. Constellation is estimated to be composed of 648 to 882 communication satellites in 18 polar orbits at 1,200 km height [4]. This is the first mission in LEO that will use Ka and Ku band for communications. Thus, specific requirements are generated in this mission in order to avoid interferences with GEO satellites using these bands. The solution is based on slightly spinning or tilting the satellite emission direction when passing through equatorial latitudes. This solution is called Progressive Pitch and is a patent of 7

20 Introduction OneWeb. A deeper explanation of that solution can be found in OneWeb website [4] and clarifying representations in Figure 1-5 OneWeb satellite [4] Services will be provided to final users by OneWeb user terminals communicating with the satellite. Each satellite will have the opportunity to connect with 50 to 70 ground stations called gateways. Each satellite will be 150 kg size capable of delivering 7.5 Gbps with GHz in gateway downlink, GHz in gateway uplink, GHz in user downlink, and & GHz in user uplink [4]. 8

21 Introduction Figure 1-6 Interferences with GEO satellites Figure 1-7 Progressive pitching used for avoiding interferences with GEO satellites 1.2 Mission Statement The proposal will be charged with providing a continuous global coverage for internet or specific telecommunication service demanded by the customer, from a Low Earth Orbit constellation. Customers in each country will be local telecommunication companies which in turn provide services to private users, government agencies or emergency and military organizations, thus making them able to compete with other satellite communications providers and expand their business more easily. The overarching goal is to generate profit for us and our 9

22 Introduction telecommunication company customers through the use of innovative solutions, based on a flexible constellation of small satellites. The target fully operative date will be 2025, in order to compete in the market with competitors such as OneWeb. 1.3 Mission Baseline Main Mission Parameters Spacecraft mass 150 kg Spacecraft power 290 W Spacecraft lifetime 8 years Global coverage Yes Continuous coverage Yes Number of satellites 245 Orbit altitude 2,000 km Main spacecraft mass and power values is based on initial budgets estimations, explained in section 1.5.1; lifetime is based on cost and revenue forecast to make the mission profitable; and orbit parameters based on constellation design. The designed mission constellation started with a walker delta feasibility study in order to distribute the condensed polar coverage. After evaluation, polar orbits showed to be more efficient and are the final choice. Moreover, this is the main choice of developing and developed global coverage missions. These solutions qualify for global and continuous coverage in a simple way. Satellite design process was based on get the highest data rate possible in a 150 kg satellite. Ground stations allocation on ground is not considered in this study. Final constellation design parameters are 11 polar orbit planes with 22 satellites per plane. It makes a total of 242 satellites. However, for reliability issues, 3 spare spacecraft will be delivered fulfilling all the launcher weight capabilities. Final orbit height is 2,000 km since total mass of the mission is significantly reduced with height and this is the maximum permitted height to comply with client requirements to remain in LEO. It will require more propellant per launch, but the principal launching cost factor is the total mass delivered. 10

23 Introduction Communications will use Ka band for ground station communications at 25.3 GHz, K band for inter-satellite communications at 22.5 GHz, and S band for telemetry tracking and command at 2.4 GHz. Two antennas are used to generate an omnidirectional communication system for TT&C, two of the K band antennas are used for inter-satellite communication along the same orbit and the other two for communication with satellites in adjacent orbits, and two antennas are used for main downlink and uplink for user communications. Regarding the data rates, Ka band links are capable of delivering 155 Mbps, K bands 50 Mbps and S bands 10 Mbps. In this mission, since it is using Ka bands for ground communications, will also interfere with GEO satellites. This mission implements an array phased antenna to change the direction of the emission and avoid these interferences. In addition, this antennas technology enables the satellite to focus on covering more demanded areas on Earth, in terms of communications. Figure 1-8 satellite design Figure above shows a simple and not fully representative model of the satellite. In the bottom is shown the phased antenna for user communications. Solar array is deployed in the opposite direction of the main propulsion system used for orbit housekeeping to avoid damage on the power system due to the expelled gases. 11

24 Introduction However, its position affects the moment of inertia and centre of mass, and must be properly studied together with thrusters positions. Internally, tanks, reaction wheels and electronic components are distributed as much symmetrically as possible to minimize perturbing torques. The other subsystems are top-level described below: The power subsystems are capable of generating 321 W EOL. It applies GaAs NeXt Triple Junction (XTJ) Prime Solar Cells and Rechargeable Li- Ion. AOCS and Propulsion systems will share their propellant tanks to minimize the total mass of the system. It imposes that both will use Mono-methylhydrazine (MMH) and N2O4 as propellants. Six bipropellant thrusters will be applied in total. Two of them will be used for orbital manoeuvres and the other four for the AOCS system. In total, 33.4 kg of fuel will be stored in the satellite for orbit and attitude manoeuvres ate the beginning of the mission. The thermal control system will utilise passive systems, optical solar reflectors and cover white paint. To conclude, the structure will be manufactured using aluminium honeycomb of 25 mm width based on actual designs of SSTL small satellite blueprints. Cubic shape is chosen for better exploiting launcher available cargo volume. 1.4 Mission Requirements Following mission requirements are based on previous mission baseline description and specific subsystem studies. Functional The small communications satellite shall be capable of delivering 50 Mbps of data connectivity from the LEO The mission shall provide continuous global coverage The mission shall provide inter-satellite communications 12

25 Introduction The satellite shall be able to maintain their orbital station The satellite shall be able to close the communication link to small antennas in the ground Lifetime shall be estimated so that the mission is profitable The concept shall present at least an innovative concept or solution Reliability. TRL shall be not less than 7. The satellite shall be flexible enough to cope with different customer needs. Satellite lifetime will be 8 years based on cost estimations. Operational The mission shall provide a method of safely disposing the satellite at the end of the mission life Satellites conforming the constellation shall have access to a ground station network for telemetry and command purposes Mission design shall maximise as much as possible service profit Constraints The weight of the satellite shall not exceed 150 kilograms The satellites shall not interfere with satellites in the GEO Constellation and global coverage shall be available by 2025 OneWeb is our main competitor. They plan to have the constellation and service available by Drivers The weight of the satellite shall not exceed 150 kilograms Constellation and global coverage shall be available by Budgets In this section is justified the initial estimations and explained the main changes compared to the final design. 13

26 Introduction Initial Estimations These values are based on Wertz, et. al. [5] mass percentages breakdowns considering that the total mass with 15% margin is 150 kg. Table 1-1 Initial mass budget Subsystem Percentage (%) Total Mass (kg) Margin (%) Final Mass (kg) Payload/Communications % 46.5 Structure % 40.5 Thermal Control % 3.0 Power % 31.5 TT&C % 3.0 OBDH % 7.5 AOCS % 9.0 Propulsion % 4.5 Other (balance + launch) % 4.5 Total % To calculate the initial power budget, the following required power breakdown was obtained from Jin, et. al. [6], who present a power budget for a small communication s satellite. Required total power was scaled using a linear relation between mass and power. Table 1-2 Initial power budget Subsystem Percentage (%) Total Power (W) Margin (%) Final Power (W) Payload/Communications % Structure % 2.9 Thermal Control % 29.3 Power (including harness) % 26.4 TT&C % 35.2 OBDH % 35.2 AOCS % 29.3 Total %

27 Introduction Final Budgets and Comparison Final mass budget is presented below. This table was generated after the preliminary design of each subsystem. Subsystems breakdown has been reorganised: Launcher systems mass has been introduced in structure subsystems. Table 1-3 Final mass budget Subsystem Percentage (%) Total Mass (kg) Margin (%) Final Mass (kg) Payload/Communications 8% % 11.3 Structure 11% % 16.4 Thermal Control 1% % 1.1 Power 17% % 25.3 TT&C 1% % 1.1 OBDH 4% % 5.7 AOCS 15% % 22.6 Propulsion 44% % 66.5 Total 100% % Figure 1-9 Final mass budget percentages In order to qualify for mission mass requirements, the margin available in the system after the first design is reduced from 15% to 13%. Below are presented the main changes in subsystem masses. 15

28 Introduction Figure 1-10 System mass comparison between initial estimation and preliminary design mass As can be seen in the graph above, propulsion and AOCS where the main subsystems that present greater differences respect to the initial estimations. This is mainly because the constellation was chosen to be placed in a very high orbit with their associated transfer orbits at BOL and EOL, and the avoidance manoeuvres, since operational orbit is at 2,000 km height. Structure subsystem also presented a big difference respect to the initial estimation, but in that case because the utilised materials are lighter than in former missions. Finally, payload and/or communications presented a reduced mass budget since the available mass was significantly reduced due to the required propulsion capabilities. 16

29 Cost Analysis 2 Cost Analysis In order for the first draft of a space mission to be valuable, there is a need for a fast assessment of its economic feasibility, as well as the technical readiness of the industry that is needed to carry it out. This has to be taken into consideration before a bill of materials is available. This section contains cost estimations performed in this project. First one is based on initial mass budget estimation and was used to calculate the lifetime of the satellites. The last one is a proper cost estimation using preliminary design mass values. 2.1 Estimating Mission Lifetime A possibility for first estimates of the costs involved in a space mission can be obtained by simply propagating the costs of reasonably similar, past space missions. Let us consider the mission to be similar to that of surveillance or meteorological satellites, which have an estimated historical cost of k$/kg [7]. Accounting for inflation, this would give us a specific cost range of k$/kg in 2017 US dollars. Taking the mission requirement of maximum 150 kg, and a middle point of 132 k$/kg, the cost per spacecraft would be 19.8 million dollars. Costs, however, may be decreased through the use of economies of scale and learning factors during the manufacturing of the satellites, so let us consider a cost of 100 k$/kg to account for the fact that the constellation involves a lot of satellites, and thus a lot of opportunities for lessons to be learned. In order to reach this costs figures, a high degree of standardisation will need to be achieved in the manufacturing processes and the launch operations. The implementation of COTS components and multiple-deployment launces are also key to achieving the target cost reductions The cost breakdown for a typical mission as provided by NASA is given in the next figure: 17

30 Cost Analysis Figure 2-1 Average NASA small spacecraft mission With these figures, 800 satellites may be launched with a cost of 12 billion dollars. In order for the mission to be profitable, the revenue must obviously exceed the cost, but historically, demand estimates have been the biggest problem to the success of this type of missions. A Cost Per Function model is proposed to estimate the price to be charged to private customers willing to use the system: T k I (1 + CPF = 100 ) T + i=1 C ops,i T C s L f,i i=1 With I being the total investment cost, k the interest rate over T years of life, C ops,i the operating costs for year I, C s the number of channels the system can support simultaneously, and L f,i is defined by: 18

31 Cost Analysis N u A u L f = min { C s 1 With N u being the estimated number of users and A u being the average user activity expressed in minutes per year. With current day communications, a more precise measure of user activity may be the amount of mobile data they use. According to Ericsson [8], the average smartphone user in 2021 will use 8.9 GB of data per month, and it is expected to grow. Fully providing this amount of data with our expected data rate would amount to seconds of system operation per month, or minutes per year. The same report also estimates a total worldwide mobile subscription number of 9 billion by Assuming deals are in place with local providers to back up their ground-based services, a market share equivalent to 0.5% of this figure could potentially be achieved. Substituting in the formulae above for a system able to support channels globally: Or in a case with 0.1% of market share: L f = L f = 0.05 From the cost breakdown, the equation for a break-even lifetime is: CPF = I (1 + k 100 ) T + T C ops,i T C s L f,i Let us assume an accrued interest rate of 5% and cost operations equal to 8% of total costs as per mission breakdown. 19

32 Cost Analysis [(1 + k T 100 ) ] CPF = T L f Though a typical modern-day tariff is close to a price of 1$ per 100MB, these prices are likely to decrease, probably bringing them closer to 0.5$ per 100MB or even lower by 2025, the intended launch date. If we assume again that we can get 50% of this price after provider margins and taxes, and 100MB are equivalent to 16 seconds of system operation, the system operation may cost 0.94$/min. With the more optimistic estimate, the system would break even by its second year of operation, the more pessimistic one ends up being profitable after its 7 th year. The proposed lifetime of the mission is to be then 8 years, as it seems a reasonable figure for profitability and technological obsolescence. 2.2 Final Cost Estimation In order to provide a first order of magnitude estimate, the most common procedure is to propagate historical data on previous space missions, preferably some which are similar to the one being proposed. That is of course, if it is in fact impossible to obtain data corresponding to the existing mission plan. The model used was Aerospace Corporation s Small Satellite Cost Model (SSCM), which uses subsystem s estimated weight as inputs. The reason to use weight as a parameter is that it has been proven a reliable predictor of mission cost. An additional factor of 1.3 was added to the cost of the payload due to technology readiness level considerations. TLR for this type of electronically steerable antenna was estimated as 5, because the research done shows that, whilst it is widely used for radar applications, it has not been flight proven for communications purposes. 20

33 Cost Analysis This model has been created to estimate the costs of small spacecraft up to 400 kg of weight, and it considers development and construction costs up to the first flight unit. After obtaining these costs, we will assume that every subsequent unit can be built for an approximately 90% of the same cost, and the development costs are distributed among them. This is a rather conservative estimate, but this fact is partially compensated by the 13% mass margin in each subsystem. For the final constellation of 245 satellites, including spares, the cost per satellite after the first protoflight unit comes close to 20 million dollars accounting for inflation. Large improvements on this figure could still be made on this figure during the development process, so if the mission can be profitable (as estimated in section 2.1) with this estimate, it is likely to produce a better financial case. Software is assumed to be the same for all satellites in the constellation, so its cost is only attributed to the first unit too. From Space Mission Engineering: The new SMAD [5], the software cost may be estimated as 500$ FY2010 per Source Line of Code (SLOC). Previous reference for space communications missions from the same book confirms that SLOC for the total system may be a reasonable estimate. This is again a very conservative estimate, as most projects are below this SLOC requirements according to the same book. The lack of available data on operations costs makes it necessary to use another tool to estimate them. The NASA small satellite mission cost breakdown (Figure 2-1) will be assumed to be a reasonable tool to define them, so the available cost estimate will be added an additional 8% cost to account for these. Cost estimates for landing rights could not be found, but they are expected to be negligible compared to the overall program cost. Ground support equipment development and construction are attributed to the first flight unit. Table 2-1 presents a final overall mission cost around 6.3 billion dollars. Presented values are updated to fiscal year

34 Cost Analysis Table 2-1 Estimated cost of the mission Concept First unit cost / k$ Structure 1,296 1,165 Thermal control ADCS 2,291 2,059 EPS 7,065 6,351 Propulsion TT&C OBDH 1,477 1,327 Payload 7,022 6,312 Integration Assembly 1,877 1,687 & Testing Software 24,670 0 Program management 3,092 0 Subsequent unit cost / k$ Global costs / k$ Launch 863,000 Ground support 891 Total per S/C 51,057 20,138 4,964,661 Operations 466,213 Total 6,293,874 22

35 Satellite Constellation 3 Satellite Constellation In order to meet the requirements, set for eventual global coverage, the optimal constellation is the polar constellation. This enables the constellation to achieve seamless global coverage that is required for any service provider (communication, internet etc). Figure 3-1 Final constellation In order to stay within the Low Earth Orbit (LEO) limit as well as to cover as much of the globe as possible with each satellite, the chosen orbit altitude for the constellation is at 2000km. The active constellation has 11 orbital planes, 22 satellites per plane, 242 satellites in operation at any given moment. Each orbital plane would also house two spare satellites each at lower altitudes, bringing the total number of satellites in orbit up to

36 Satellite Constellation 3.1 Constellation Requirements and Constraints The constellation for the mission is required to provide eventual global coverage for providing data services. As such the completed constellation has to cover all sections of the globe reliably at all times. At the same time given that global data transfer and communication is also a part of the mission, one of the constraints is that the satellites are required to be in view of each other, enabling intersatellite communication within the constellation. This will be further discussed in the design section of the report. Given the constraint of the mission to be in LEO, and the satellite mass limited to a maximum of 150 kg (which restricts payload size and power), it can be initially assumed that the number of satellites required to achieve global coverage will be significantly large. This means that the full constellation is also going to take a significant amount of time to be completed. Considering the operational lifetime of each of the satellites, it is important to start to utilise them as soon as possible to make the most of the time the spacecraft spends in orbit. Thus, a minimum constellation configuration is advisable. This minimum configuration will not be able to provide seamless coverage as there will be gaps within the constellation. Thus, feasibility of the potential minimum configurations should be evaluated in terms of service interruption. The final consideration is that, though the service has to be global at the end of the mission, the majority of the world s population resides with ±60 latitude. Thus, statistically speaking, most of the service users should also be situated within this region. Potential constellation geometries were considered whilst keeping this factor in mind. 3.2 Initial Concepts & Considerations There are a limited number of standard constellation geometries available for implementation within a mission such as this. As it is a global service providing mission requiring inter-satellite communication, the satellites in each plane will have to remain within observable distance of each other throughout its orbit. By 24

37 Satellite Constellation doing so, they are able to communicate to each other within the same plane. Communication to satellites in other planes can be carried out nearer to or at the orbital converging points as demonstrated below. Figure 3-2 Orbit plane intersections Considering these factors, the potential geometries are the Walker Delta Constellation with additional polar orbit planes and the standard polar orbit constellation. A polar constellation is able to cover the surface area of the entire globe whilst maintaining seamless, continuous coverage. Maintaining line of sight with satellites in plane and adjacent plane is also easy in a polar constellation. This is due to the fact that all the satellites in the constellation are affected identically by the perturbations caused by the Earth s Oblateness (J2 Perturbations). Thus, in theory, the satellites need only be oriented as required at the beginning of the mission. For the rest of the mission lifetime, they can remain as they are. 25

38 Satellite Constellation Figure 3-3 Polar constellation A walker delta constellation at 60 is able to cover the heavily populate zones on the earth. However, the Polar Regions are left without any coverage whatsoever. This can be remedied with further orbit planes in the Polar region with further satellites in place. Alternatively, this can be addressed by re-directing the beams of the satellites at higher latitudes towards the polar region. This is possible due to the fact that in a walker delta, at the latitude of the chosen constellation inclination (i.e. 50 latitude for a 50 inclination constellation), there is overlaps of multiple satellites within that region. For example, in Figure 3-6, only two of the planes are visible. This enables us to observe that there are two satellites near the intersecting points of the orbit. As only one satellite is required to maintain ground contact at that point, the other satellite can be re-oriented to target the Polar Regions. As there are further planes in the walker delta spread around the globe, a full coverage can be achieved. 26

39 Satellite Constellation Figure 3-4 Polar targeting However, a major issue remains with the Walker Delta, which is that with identical number of satellites to that of the Polar constellation, there remain gaps within the constellation, as shown in Figure 3-5 and Figure

40 Satellite Constellation Figure Walker delta with square beams Figure Walker delta with circular beams 28

41 Satellite Constellation This issue can be addressed by changing the beam shapes to geometrically cover as much of the surface area as possible. As it can be seen, that by changing from the square beam to circular improves the coverage. 3.3 Design of Final Constellation Though other more novel constellation designs can be implemented to achieve global coverage, the number of satellites required for global coverage at LEO is considerably large. Thus, in order to reduce mission complications whilst achieving continuous global coverage the final decision is to opt for the Polar constellation. Though the Walker Delta can be utilised for this mission after further parameters are implemented into the mission (changing beam shape, changing, orientation etc), this adds further complications to the mission that are unnecessary. As the mission is already quite involved, it is better to keep most of the design simple. In order to maximize the footprint of the beams on the ground the orbit altitude is set at 2000km. Table 3-1: Satellites and planes required at different altitudes Altitude Satellites per plane No. Planes Total Satellites

42 Satellite Constellation Number of satellites required at each altitude can also be calculated roughly using trigonometry (altitude and footprint size). As a slight overlap is required for seamless coverage, the number of satellites comes down slightly the given value in Table 3-1. Thus, at our chosen altitude, the required number of satellites comes down 242 as opposed to 248. The final consideration is the minimum configuration required for service inauguration. Given the choice of polar constellation, any form of useable mission structure can only be achieved when at least 5 planes have been achieved for the mission. As the chosen Launch Vehicle (LV) is capable of launching 60 satellites at a time, within two launches a minimum configuration can be achieved. However, some thought needs to be given to what the required geometry should be for the minimum configuration. By spacing out the 5 planes as equally as possible (whilst considering the completed geometry), service disruption can be minimised to ~1hour at the equatorial region (where the disruption occurs for the longest period of time). This means however that the LV will require higher Delta V and subsequently more fuel to carry out the necessary plane changes. Leaving very little margin for error this will increase the required number of launches and thus time needed for the full constellation to be achieved. 30

43 Satellite Constellation Figure 3-7 Minimum configuration with 36 RAAN spacing Figure 3-8 Minimum Configuration with 18 RAAN spacing A better solution is to opt for the adjacent placement of the planes for minimum configuration. This will increase service disruptions to 6 hours as opposed 1 hour, 31

44 Satellite Constellation however it ll mean that the overall constellation can be achieved sooner due to reduced number of launches. Considering the total number of satellites required (including spares) and the number of satellites the LV is able to launch at any given time, the entire constellation can be completed in 5 launches, whilst the minimum configuration can be achieved in two launches. As the chosen launch vehicle is the Atlas, which launches around 8-10 times per year, the full constellation can be completed in 4-6 months and minimum configuration in as little as a month. 3.4 Requirements and Constraints on Constellation, Including Management of Interference Interference Mitigation As the constellation is located in LEO and broadcasting information down to ground using Ka bands, there will be interference with telecommunication satellite using the same band but located in GEO. 2 ways of mitigating those interferences have been planned. The first one is coordinating the ITU (International Telecommunication Union) to have specific frequencies allocated to the constellation (see section 5.1). By using dedicated frequencies, interferences should be avoided. This is the goal of the ITU and its spectrum management. The second way of mitigating interference is to move the signal in a way where it is not parallel anymore to the signal of GEO satellites. The selected way of doing that is to use phase array antennas. In a phased array antenna, the beam can electronically be steered in the wanted direction. This method will allow to only move the beam and not the entire antenna nor the whole spacecraft when crossing the GEO satellites beam [9]. These antennas are compound of small emitters that controlling their order or process of emission can concentrate the broadcasted signal power in certain 32

45 Satellite Constellation covered areas or change the direction of emission. Basically, individual antennas are barely delayed one respect each other to generate that effect. 33

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47 Communication Subsystem 4 Communication Subsystem 4.1 Introduction Every spacecraft does need in one way or the other some form of communication system. The necessity for a communication system is because of the need for telemetry, tracking and command (TT&C): for operations, or sending mission data down to the ground (user), in most cases both forms are required [5]. A basic communications system is made up of three sections: the space section which is the S/C, the control station (TT&C) and the ground or mission operations. The space section consists of onboard hardware such as: transmitters, receivers, antennas, etc. The control section consists of TT&C facilities. The ground section consists of high gain antennas for both reception and transmission [10]. 4.2 Communications Links Architecture Communications architecture is topology of communications links, it is the structure of uplinks, downlinks, intersatellite links, TT&C links and ground station. Figure 4-1 Satellite communication link architecture [10] 35

48 Communication Subsystem The communications architecture of this mission is designed in such a way that it has the following: 1. Intersatellite links 2. Ground gateways stations 3. Operations ground stations (through TT&C link) 4. Voice communications, tracking and broadband access to the following users: a. Maritime b. Land c. Aviation The designed is target every possible customer in the government, public safety, commercial, military, personal and maritime. The general architecture will look like the network in Figure 4-2. Figure 4-2 Classical satellite communication system [11] 36

49 Communication Subsystem Inter-Satellite Links To sustain a global coverage there is a need for intersatellite cross links between satellite in same and adjacent planes with a data rate of 16 Mbps; making a star network (link), these links will be RF links in K-band in the range GHz [10]. Figure 4-3 shows the intersatellite link between the planes. Figure 4-3 Intersatellite link between the planes [12] Gateway Stations As part of getting a global coverage there is a need for gateway stations to connect customers with access to services. There will be two gateway stations along each plane to ensure a continuous coverage and access Communication Payload Repeaters The communications payload consists of the repeaters and antennas, the satellite repeaters are regenerative: unlike transparent repeaters they provide improved link quality. On-board processing (modulation and demodulation). 37

50 Communication Subsystem Figure 4-4 Transparent and regenerative repeaters [10] Antennas The mission will constitute: 1. Four K-band phased array antenna antennas on each spacecraft for intersatellite links 2. One Ka-band phased array antenna for transmission 3. One Ka-band phased array antenna for reception 4. Two S-band antennas for TT&C to provide omnidirectional illumination Receivers and Transceivers There is a need for receivers and transceivers to support the transmissions via the antennas, the following are the chosen configuration: 1. Four K-band transceivers 2. One Ka-band receiver and transmitter each 3. One S-band transceiver 4.3 Link Budget Link budget is the accountability of all losses involved in a communication link, Figure 4-5 shows the idea clearly. 38

51 Communication Subsystem Figure 4-5 Link budget picturing [13] Link budget is governed by the equation: E b = PL sl l G t G r N 0 k B T s R Where: P = transmitter power (W) L s = free space loss L l = other losses G t = transmitter gain G r = receiver gain k B = Boltzmann constant (= 1.38 x WK 1 Hz 1 ) T s = system noise temperature (K) R = data rate (bits per second) 39

52 Communication Subsystem Frequency The frequency used: K-band: 22.5 GHz Ka-band: 25.3 GHz S-band: 2.4 GHz Data Rate The date rate of the frequencies employed are: K-band: 50 Mbps Ka-band: 155 Mbps S-band: 10 Mbps Orbit The orbit altitude is 2000 km; therefore, the maximum time of view is seen as in Figure 4-6. Figure 4-6 Spacecraft line sight geometry [14] At 2000 km altitude and 10 elevation, the following parameters are: Range max. path length: 4437 km 40

53 Communication Subsystem Maximum time in view: approx. 22 mins Table 4-1 Link budget calculations Parameter K-band Ka-band S-band Frequency (GHz) Data Rate (Mbps) Altitude (km) Max. Path Length (km) Propagation Loss (db) Transmitter Power (W)

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55 Regulation Aspect and Interferences Mitigation 5 Regulation Aspect and Interferences Mitigation The global coverage of the constellation intends on providing service internationally. This brings about concerns onto the regulation aspect of satellite telecommunication: can we broadcast in another country than our own? What do we have to do to be allowed to? This section aims at answering those questions. 5.1 Landing Rights and Spectrum Management There are several types of license needed to operate a satellite telecommunication system and provide service. Satellite operator needs to get a space segment license to be able to operate. Licensing is mandatory to regulate limited resources such as the frequencies used and the orbital slot allocated (especially in GEO). Landing rights refers to the need to ask authorisation to broadcast in a country. While this still exists in several countries where it is necessary to ask for those authorisations, the open sky policy renders those useless in most countries. The ITU (International Telecommunication Union) is in charge of managing the radio spectrum and obits allocation at an international level. This Union was created out of a need to avoid interference when 2 satellites uses the same frequency for example. With the open sky policy, once a satellite operator was granted a license to use a part of the spectrum at the ITU level, there is no further need of a license in another country [15,16]. Frequencies allocation is done by the government of the country where the satellite operator is registered. As part of their spectrum management program in line with the ITU, each country has a Frequency Allocation Table (FAT) that they use to grant a license to a satellite operator [15,17]. Those tables describe for each frequency available, the service that can be done, the region that can be covered and other specifications. Services can be: fixed (FSS, fixed satellite service), from the satellite to a fixed station on the ground, this is the service that we provide, 43

56 Regulation Aspect and Interferences Mitigation mobile (MSS, mobile satellite service), from the satellite to a mobile station on the ground, i.e. satellite phones, broadcasting, space research and Earth observation, radiolocation, radionavigation, etc. Regions are defined by the ITU: region 1 is Africa, Europe, Middle East and Russia, region 2 is the American continent, region 3 is Asia and Oceania. The UK FAT has available frequencies in the Ku and Ka bands for FSS covering all 3 regions. The ground segment (here the small antennas) also requires licensing. The entire licensing process both for the ground and space segments takes time and has a cost. However, information on the delay, the cost and the ground license where not readily available. 44

57 Propulsion Subsystem 6 Propulsion Subsystem This chapter of the report describes the propulsion subsystem design considered for the communication constellation proposed. The initial section lists the options considered and their parameters as well as it defines the approach taken in designing the subsystem. Followed by that is a description of the fuel storage subsystem and the steps taken into sizing it, defining its parameters and characteristics. A summary of the subsystem with a graphical representation concludes this chapter of the report. 6.1 Engine Selection Options Space engines and thrusters can be characterised by many different parameters. However, the most important ones are: Engine type, i.e. what method is used to provide thrust. Engine types vary from solid and liquid propellant engines, through Ion and Hall Effect thrusts to resistojets, arcjets, solar sails, etc. Specific Impulse (ISP) which is a measure of the engine efficiency. High ISP corresponds to more efficient engines and vice versa. Thrust defining the thrust an engine can provide. High thrust engines are preferred for LEO space mission due to that quicker acceleration reduces gravity losses and thus increasing manoeuvring efficiency. Engine mass and complexity. Power requirement either for the operation of valves, turbo pumps and mechanical devices or for molecular acceleration in the case of Ion thrusters. For the design of the communication constellation proposed, multiple engines different in design and thrust generation method are considered. Generating a list of different options is key for identifying the most feasible and most costeffective solution. Table 6-1 below, provides a list of the thrusters considered for 45

58 Propulsion Subsystem the mission designed. Engine thrust and mass, specific impulse as well as engine type are listed for each of the options allowing comparison between the different engines considered. Table 6-1 Thrusters considered Engine Manufacturer Thrus t (N) s ISP m/s Engine Mass (kg) Type NSTAR Min NASA ,700 16, Ion NSTAR Max NASA ,100 30, Ion NEXT NASA ,100 40, Ion T5 Min Qinetiq ,500 34, Ion T5 Max Qinetiq ,500 34, Ion T6 Min Qinetiq ,400 43, Ion T6 Max Qinetiq ,400 43, Ion BHT-8000 Busek ,210 21, Hall Effect BHT-1500 Min Busek ,615 15, Hall Effect BHT-1500 Max Busek ,865 18, Hall Effect LEROS-1C Moog ISP ,188 4 Bipropellant DST-11H Moog ISP , Bipropellant TR-308 Northrop Grumman , Bipropellant TR MN Northrop Grumman ,188 6 Bipropellant TR YN Northrop Grumman ,237 6 Bipropellant R-4D Aerojet , Bipropellant Model S Airbus , Bipropellant 200 N Airbus , Bipropellant MRE-1.5 Northrop Monopropellan , Grumman t ARC-445 Aerojet , Monopropellan t MR-502A Aerojet , Resistojet MR-509 Aerojet , Arcjet At an early stage of mission design, it was agreed that liquid propulsion is the most feasible option for the satellites considered as the other solutions do not fit 46

59 Propulsion Subsystem with mission requirements. For example, Ion propulsion demands high power performance which is not achievable by a small spacecraft. In addition to that, the low thrust characteristics would increase deployment time and would increase burning time due to gravity losses. Comparing liquid and mono propellant engines, the two feasible options were identified. The first being the Airbus manufactured 200 N thruster and the second being the MRE-1.5 designed and manufactured by Northrop Grumman. The higher specific impulse of the bipropellant option meant that lower fuel mass would be required while the monopropellant option would provide simplicity and dry mass savings due to the lighter structure of the thruster and its related components. For this reason, a study was conducted to select the more feasible option of the two. The approach implemented included taking into account the thrusts (2 thrusters for safety reasons) dry mass and the mass of the fuel required for mission execution and then comparing the results. It must be noted that due to the robust computational model created, different deorbiting profiles (end of life burn) are considered. This is to comply with the requirement of complete deorbiting within 25 years of mission finish as described in Chapter 8 of this report. Table 6-2 MRE-1.5 & 200 N efficiency comparison Engine & Fuel Mass including 20% fuel margin MRE N Difference in overall mass As it can be seen from the results obtained, the 200 N bipropellant thruster manufactured by Airbus is a more feasible option as it saves about 2.7 kg of overall weight. For this reason, each satellite in the constellation will be equipped with two 200 N engine ensuring efficiency and reduced risk in the case of engine failure. The introduction of this type of engine to the spacecraft design simplifies the overall orbiter design due to that the ADCS thrusters can be implemented to the propulsion subsystem as ADCS thrusters use identical fuel and design. This 47

60 Propulsion Subsystem significantly reduces design and manufacturing issues as well as it is a more costeffective solution due to shared fuel tanks Airbus 200 N Bipropellant Thruster Airbus s 200 N bipropellant engine has been originally designed as an attitude control, manoeuvring and braking thruster for ESA s Automated Transfer Vehicle. However, due to its lightweight and relatively high efficiency the engine is a great fit for the constellation developed. Another advantage of the engine is its reliability as a result of the amount of missions flown. Table 6-3 Airbus 200 N Bipropellant thruster technical specifications states the key technical parameters of the 200 N bipropellant thruster while Figure 6-1 provides a sketch of the engine selected. Table 6-3 Airbus 200 N Bipropellant thruster technical specifications Airbus 200 N Bipropellant Thruster Technical Specifications Nominal Thrust 216N ± 10N Thrust Range 180N ± 15N to 270N ± 15N Specific Impulse at Nominal Point > 2650 Ns/kg (> 270 s) Fuel MMH Oxidizer N204 Fuel Density g/cm 3 Oxidizer Density g/cm 3 Mixture Ratio Nominal 1.65 ± Mixture Ratio Range 1.2 to 1.9 Flow Rate Nominal 78 g/s Flow Rate Range 60 to 100 g/s Injector type Impingement with film cooling Chamber Pressure Nominal 8 bar Inlet Pressure Range 17 ± 7 bar Nozzle End Diameter (inner) 95 mm Throat Diameter (inner) 12 mm Nozzle Expansion Ratio (by area) 50 Chamber / Nozzle Material SiCrFe coated niobium alloy Minimum on time 28 ms Minimum off time 28 ms Maximum burn time (single burn) 7600 s Number of full thermal cycles 250 Valve power requirement 32 W Total mass 1.9 kg 48

61 Propulsion Subsystem 6.2 Fuel Storage Figure 6-1 Airbus 200 N Bipropellant thruster sketch The introduction of the 200 N space thruster to the design of the orbiters significantly simplifies the fuel storage subsystem due to the fuels used in combustion and the mixture ratio of those fuels. Combustion of Monomethyl hydrazine (MMH) and Dinitrogen tetroxide (N204) at a ratio of 1.65 defines that the fuel tanks for the oxidiser and the fuel tanks for the fuel are of identical size. Due to that, sizing and packaging of the fuel storage subsystem is significantly optimised. Fuel tanks selection is based on comparison of different pressurised fuel tanks, space certified for MMH and N204. This is followed by computing their storage capabilities and identifying the number of tanks required to carry the mission fuel which is 29.9 kg in total. Three of the fuel tanks considered proved to be adaptable to the spacecraft design. All those three fuel tanks are manufactured by MOOG-ISP in the United States of America. Table 6-4 shows the computation performed for each of the final options considered. 49

62 Propulsion Subsystem Table 6-4 MOOG-ISP fuel tanks considered Fuel Tank Volume (cm³) Mass per tank (kg) Diameter (cm) Tanks needed Oxidizer Fuel Tanks mass (kg) Hit to Kill Exploration Target As it can be seen, the Exploration fuel tank is too large for the mission design. This would mean that extremely high volume of helium should be carried on for pressurisation as well as the empty weight would be significantly higher in comparison with the other two options. The other two options show to be capable of carrying the fuel required for mission execution. Hit to Kill option would require two fuel and two oxidizer fuel tanks (4 in total) while the Target fuel tank design would require a fuel tank for each of the fuel and oxidiser. Also, both options would be launched fuelled by over 90 % (98% for Hit to Kill and 93% for Target fuel tank) meaning empty mass is reduced to minimum. However, due to the lighter overall weight, the Hit to Kill option is identified as the most feasible option for the orbiter s fuel storage systems. 6.3 Propulsion Subsystem Summary In summary, each of the satellites launched will be equipped with two 200 N bipropellant thrusters as main engines. The engines will be fed by four fuel tanks two containing Monomethyl hydrazine (MMH) and two containing Dinitrogen tetroxide (N204). Helium will be used for fuel tank pressurization as fuel is drained. The fuel tanks will also feed the four ADCS thrusters. Figure 6-2 shows a schematic representation of the propulsion subsystem of each of the satellites. Pressure sensors, drain valves, regulators, filters, non-return safety valves and check valves are introduced throughout the system to reduce risk and provide control over the subsystem. 50

63 Propulsion Subsystem Figure 6-2 Propulsion subsystem design The overall mass of the propulsion subsystem including fuel (plus 20% margin), pipes, storage tanks (including helium) and main engines but excluding ADCS thrusters is 40.7 kg which is about 27% of the spacecraft wet mass. 51

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65 Launch and Orbit 7 Launch and Orbit This chapter of the report describes the process implemented in launcher selection as well as a brief summary of the launch events for the communication mission designed. In addition to that, the deployment procedure of constellation is discussed together with, delta V and propellant mass budgets. 7.1 Launcher Selection Selection Method Launch vehicles come in wide range of performance and capabilities. Many launch vehicles are designed for a specific market, for example GTO or LEO. Low Earth Orbit launchers have a significantly wider range due to the bigger market and the ease of access in comparison to GTO or interplanetary launches. The key parameter defining launcher selection is the final orbit at which the launcher releases its payload (spacecraft). Based on the constellation requirements, 242 satellites orbiting in eleven polar planes at a circular altitude of 2000 km are required for the smooth operations of the service. However, releasing the payload at the exact final orbit would be a complex procedure requiring multiple manoeuvres and a long-time span to separate the spacecraft in order to provide global coverage. For this reason, early on a decision was made to launch the satellites into a lower orbit. Once released from the launcher, the spacecraft would be raised into the final orbit one after another. This would significantly reduce risk and the deployment time of the constellation. Four different launch orbits are considered: I. Polar 1800 x 1800 km. III. II. Polar 1800 x 1900 km. IV. Polar 1900 x 1900 km. Polar 1800 x 2000 km. Option lv was the first to be eliminated due to that it requires a more complex launch profile and because it would cross the final (operational) altitude causing risk to operational satellites. The launch profile was the reason for excluding the second option from the list. Comparing options l and lll, it was confirmed that more 53

66 Launch and Orbit mass can be launched into the lower orbit, however this comes at a penalty due to the increase in the mass of the fuel required to raise the orbit into a 2000 x 2000 km. Table 6-1provides the results obtained for propellant mass requirements for the final two options considered. Table 7-1 Propellant mass for orbital boost Polar 1800 x 1800 km Polar 1900 x 1900 km δ Propellant mass (kg) % Propellant mass (kg) As it can be seen, fuel mass increase of about 100% is required when placing a satellite into a lower orbit. For this reason, it was agreed that the satellites should be released into the high orbit of 1900 x 1900 km. Once the launch orbit is defined, multiple launcher vehicles, type capabilities and specifications are studied in order to select the most feasible launcher. The launchers considered are: Antares Atlas V Delta ll Delta lv Proton PSLV Rokot Soyuz VEGA A list with performance capabilities to the selected orbit was created for each of the launchers. Following that, comparison and trade-off was performed in order to select the most feasible option. In addition to that, different launchers and launcher combinations were assumed to reduce deployment period by as much as possible. The computations performed showed that the Atlas V 500 series would be the most feasible launch vehicle option as it would allow full deployment with six successive launches. Another important aspect taken into account is the unprecedented success rate of Atlas V launchers ensuring low risk and reduced insurance costs. 54

67 Launch and Orbit Atlas V 500 Series Parameters Atlas V is a modern spacecraft launcher operated by the United Launch Alliance (ULA) which is a joint venture between Lockheed Martin Corporation and the Boeing Company. Atlas V uses a standard Atlas Booster, zero to five strap-on solid rocket boosters (SRBs), a Centaur in either the Single-Engine Centaur (SEC) or the Dual-Engine Centaur (DEC) configuration, and one of several Payload Fairings (PLF) Invalid source specified.. A three-digit (XYZ) naming convention was developed for the Atlas V 400 and 500 series to identify its multiple configuration possibilities where X represents the diameter of the PLF, Y represents the number of solid rocket boosters and Z stands for the number of Centaur engines. The Atlas V 400 series employs the flight-proven 4-m diameter payload fairing in three discrete lengths: the Large Payload Fairing (LPF) 12.0m in length; the Extended Payload Fairing (EPF) 12.9m in length; and the Extra Extended Payload Fairing (XEPF) 13.8m in length. Similarly, the Atlas V 500 series employs three lengths of the flight-proven 5.4-m diameter payload fairing: the 5.4- m short PLF, 20.7m in length; the 5.4-m medium, 23.5m in length; and the 5.4m long, 26.5m in length. The capability of configuring the launcher first stage based on the mission parameters and variety of payload fairings, provides unprecedented flexibility to the launch capabilities of the Atlas V launchers. For the communication designed, Atlas V 500 series will be used as more satellites can be launched with a single launch. This is possible due to the relatively low wet mass of a single satellite. Also, as the parking orbit required is polar, Vandenberg Air Force Base (VAFB) launch sites will be used due to its northerly location providing easier access polar and sun-synchronous orbits. Studying the different payload fairings of the 500 series, it was calculated that 60 satellites can be launched at once using the standard short 5.4m fairing. Due to that short fairing provides maximum launch weight as no additional fairing mass is carried and due to that this fits well with the mission parameters the standard short fairing was selected for the constellation designed. In regards with payload 55

68 Launch and Orbit adapters, the Atlas V offers a wide range of adapters and separators varying in size and mass. For this reason, an average value of 120 kg is allocated for payload adapters. This mass is added to the payload mass launched at each event. Once all launcher parameters are defined, launcher performance for the selected launch orbit (1900 x 1900 km polar) are obtained for the different configurations. Table 7-2 provides a summary of the launch capabilities obtained from Atlas V Launch Services User s Guide. Table 7-2 Atlas V 500 series launch capabilities into 1900 x 1900 polar orbit Launcher Launch Site Payload Fairing Mass capability (kg) Atlas V 501 VBAF Short 5477 Atlas V 511 VBAF Short 7474 Atlas V 521 VBAF Short 9250 Atlas V 531 VBAF Short Atlas V 541 VBAF Short Atlas V 551 VBAF Short Error! Reference source not found. provides the key technical specifications of t he Atlas V 500 series Launch Procedures Ensuring launcher capabilities are maximised at every launch event is key for ensuring cost efficiency. For this reason, different launch events using different launcher specifications are simulated in order to obtain the optimal solution. Based on the results obtained, the feasible solution for placing the constellation into orbit including three spare satellites would be obtained by six launch events. The first launch will be performed on an Atlas V 521, carrying 55 operational satellites and a spare one. The next two launches will be performed by Atlas V 511 launchers and they carry 45 operational satellites plus a spare spacecraft at each of the launch events. The final three sets of communication satellites will be carried on board Atlas V 501 rockets and each event will place 33 operational satellites. This schedule would eventually place 242 operational satellites into orbit plus three spares in orbital storage. The spare mass at each of the launch events would be taken by the structure housing the satellites into their position at 56

69 Launch and Orbit launch and will allow more spares to be launched into orbit if required at a further stage of the design. The official rocket builder provided by ULA is used to provide a cost estimation for each of the launch events. Table 7-3 summarises the launch procedure implemented. Table 7-3 Constellation deployment procedure Launcher Satellites per Launch Payload (kg) 105% plus adapters (kg) Spare mass (kg) Spares Launched Cost Estimate (Million USD) Atlas V Atlas V Atlas V Atlas V Atlas V Atlas V TOTAL Insertion into final Orbit As stated in 7.1 above, the satellites will be released into 1900 x 1900 km polar orbits meaning that each of the operational satellites will need to be boosted into the higher orbit. Due to the small change in orbital altitude, standard Hohmann transfers will be performed. The transfer will be conducted by a burn raising the spacecraft apogee to 2000 km followed by a second circularisation burn at the apogee. The manoeuvres will be performed in a sequence to allow spacecraft spreading along the orbital plane. As the manoeuvres are not executed at exactly the same point, a very small RAAN spreading would occur. However, as the timespan is negligibly small, the operation of the constellation would not be affected. Thus, no additional fuel for inclination changes is added to the satellites. In the case small manoeuvres need to be carried out, 20% fuel margin is added to each of the orbiters. Figure 7-1 illustrates a sketch showing a standard Hohmann transfer as well as the key sampling points for delta-v calculations. 57

70 Launch and Orbit Figure 7-1 Hohmann transfer sampling points As stated in chapter of this report, launchers will carry orbiters for more than one of the orbital planes. Due to that satellites are released into a lower orbit in comparison to the operational orbit, a RAAN computation was performed to determine the time required for a natural switch between two orbital planes without burning propellant. The results obtained show that orbiting at a lower altitude for 57 orbits which corresponds to 5 days would result in 32.7 degrees change in RAAN. This eventually means that, the satellites orbiting in the lower orbit, would have aligned with the following orbital plane without burning any fuel or making any orbital adjustments. This period is sufficient for planning ground operations while keeping a tide schedule for constellation deployment. At the end of the mission, a third burn will be performed to lower the perigee and allow natural decay and deorbiting of the satellites. Unlike above, only one manoeuvre will be performed, i.e. no circularization burn will be conducted at the perigee. Detailed information and reasoning for this is provided in chapter 8 of this report. To summarise, two orbital manoeuvres would boost the satellites into their final orbits while a final manoeuvre at the end of their operational life would place their perigee deeper into the atmosphere resulting in natural orbital decay and satellite 58

71 Launch and Orbit burn up. As multiple satellites are launched at once, orbital planes will be deployed one after another using the effect of RAAN spreading. Natural RAAN spreading between two orbital planes would occur in about five days time which providing sufficient time for planning and execution while keeping a tide constellation deployment schedule Computation Method The theory behind the computation method used is stated below while the results are presented in Table 7-4. Calculations are performed for all three manoeuvres required orbit boos, circularization and de-orbit. Engine characteristics and the mass changes due to burning fuel are considered for every next burn. V 1 and V 4 are the circular orbit velocity speed respectively for 1900 x 1900 km and 2000 x 2000 km. V 2 and V 3 are the velocities at perigee and apogee for the 1900 x 2000 km orbit. V 5 is the perigee velocity for the orbit placing the spacecraft into an uncontrolled re-entry (details provided in section 8 of this report). Based on that, δv1 is the change in velocity placing the spacecraft into a transfer orbit while δv2 is the change in velocity required for circularisation at the operational altitude. The final δv3 value is the required change in velocity to place the spacecraft into de-orbiting mode. It must be noted that the first two manoeuvres are pro-grade while the last one is retro-grade. Velocities: V 1 = μ a ; V 4 = μ a V 2 = 2μ ( 1 r 1 1 r 1 + r 2 ) ; V 3 = 2μ ( 1 r 2 1 r 1 + r 2 ) ; V 5 = 2μ ( 1 r 2 1 r 1 + r 2 ) δv: Propellant mass: δv 1 = V 2 V 1 ; δv 2 = V 4 V 3 ; δv 3 = V 4 V 5 59

72 Launch and Orbit 1 m p = m 0 (1 exp ( δv ) ISP ) Acceleration: acceleration = T total m 0 Time to reach δv: accelaration = δv acceleration 7.3 Station Keeping and Space Debris Station keeping and debris avoidance are the two main reasons for orbital manoeuvres in a constellation. Station keeping is required due to orbital perturbations. In LEO, the main orbital perturbators are atmospheric drag, solar activity, Lunar and Earth s gravitational fields. To start with, atmospheric drag can be ignored for orbits higher than 1000 km. For this reason, it is assumed drag does not act on the constellation while in operation. The other three perturbations can also be neglected due to the high altitude and due to the relatively short operational period of the constellation. This means that any influences of those perturbators can be ignored. Moreover, any affections would act equally on all of the satellites in orbit meaning that the constellation would not be affected significantly. In regards with space debris, they vary in size and in spatial density. Of course, larger debris are more dangerous due to the higher energy they carry. Figure 7-2 show the spatial density of space debris bigger than 10 cm. 60

73 Launch and Orbit Figure 7-2 Spatial density > 10 cm (extracted from Operational Collision Avoidance by ESA Space Debris Office presentation given by Klaus Merz on 03/11/2016) As it can be seen, the probability of colliding with space debris at an altitude of 2000 km is very low. Due to the low level of risk and due to the presence of spare orbiters, debris avoidance manoeuvring can be excluded from the generic operations planning. Especially due to that, very little fuel would be required for such rare events. In summary, no additional fuel for station keeping or space debris avoidance manoeuvres is included to the propellant budgets of the satellites launched. However, it must be noted that the extra propellant carried as margin can be used for such manoeuvres if required, especially due to that such operations would be exceptional and would require minimal fuel masses. 7.4 Delta-V and Propellant Budgets Summary As discussed above, three main manoeuvres will be executed by each of the satellites launched into orbit. Two aiming to place the satellite into an operational orbit and a final one for re-entry. Due to the high orbital altitude, no extra fuel is added for orbital perturbations or debris avoidance. Table 7-4 Delta-V and propellant mass budgets shows a summary of the delta-v and propellant budgets calculated. 61

74 Launch and Orbit Table 7-4 Delta-V and propellant mass budgets Sampling Point V 1 V 2 V 3 V 4 V 5 Apogee (km) Perigee (km) Semi-major axis (km) 8, , , , , Velocity (km/s) δv (km/s) Propellant (kg) % Propellant (kg) Acceleration (m/s 2 ) Time to reach δv (sec) The results obtained show that, the circularisation burn would require less fuel and time in comparison to the first burn placing the spacecraft into an elliptical orbit even though the change in velocity is identical. This is due to the lower mass of the spacecraft at the second burn. Based in the results obtained, a total of 25.7 kg (including 20% margin) of propellant are required for orbital manoeuvring. Additional 4.2 kg (including 20% margin) would be added to the orbiters fuel tanks to serve the ADCS subsystem. In total, 29.9 kg of fuel including 20% margin would be carried on board of each satellite. 62

75 End of Mission Considerations 8 End of Mission Considerations 8.1 Disposal Options and Requirements The Inter-Agency Space Debris Coordination Committee (IDAC) is an international governmental forum for addressing the issues related to man-made space debris. The governmental body has been established in 1993 and consists of multiple space agencies and organisations governing activities in Low Earth Orbit (LEO) as well as issuing guidelines related to spacecraft/rocket stages disposal. According to IDAC, responsible removal of orbiting objects from LEO is the most major step towards limiting the growth of man-made space debris. Since 1993, the international committee studied different satellites/rocket stage disposal options and their long-term implication on space debris mitigation. Based on the results obtained and models generated, the committee concluded that limiting the amount of time each object stays in LEO significantly reduces the risk of space debris number growth. Figure 8-1 IDAC protected regions (IADC-02-01, 2007) IADC Revision 1 from September 2007, defines three disposal options for LEO satellites. The first and most preferred, is spacecraft de-orbiting at the end of the mission. This is achieved by performing an end of life burn placing the satellite s perigee less than 50 km in altitude to avoid atmospheric skip. The 63

76 End of Mission Considerations manoeuvre is also called controlled re-entry due to that, the execution must take into account the region the spacecraft re-enters Earth s atmosphere. This is performed to reduce the danger factor of spacecraft pieces surviving re-entry and causing damage to humans on Earth. IADC requires re-entry to be executed above oceans to reduce risks on population. The second option defined by IADC is disposal into a storage orbit, also called graveyard orbit. This approach requires a satellite to be placed into a Medium Earth Orbit (MEO). The document restricts placing satellites into congested MEO orbits, such as 12 hour orbits and other orbits typically used by communication and navigation satellites. Placing a spacecraft into a graveyard orbit is achieved through series of burns raising both apogee and perigee of the satellite into a safe storage region. Careful planning and execution is required to eliminate collision risks and to ensure sufficient propellant is stored for the end of life operations. This approach is useful for heavy, high altitude satellites. In the case neither of the above options is feasible for mission design, IDAC recommends uncontrolled re-entry within 25 years after the end of the mission. This is achieved by lowering the spacecraft perigee and allowing atmospheric drag to decay the orbit naturally. As orbital decay is directly related to spacecraft mass, IDAC recommends fuel tanks venting after the final manoeuvre is executed. Another reason for this recommendation is to reduce risk of explosions during the orbital decay which limits the risk of creating further space debris. In addition to that, IDAC requires detailed studies and simulations to ensure decay within the stated timeframe of 25 years or less. Natural orbital decay is widely used for Earth orbiting satellites below 800 km, for large constellations and for small satellites. While at first glance, uncontrolled re-entry might seem as the simplest disposal method, it requires detailed planning and careful execution to ensure responsible performance. Also, this approach adds operational requirements as the satellite should be monitored throughout the whole period of orbital decay to ensure decay rate, orbital position and providing warnings to population upon re-entry over populated areas. Due to increased risk to 64

77 End of Mission Considerations population, space agencies require detailed assessment of manufacturing materials to ensure all hardware will burn up upon re-entry. 8.2 Constellation Disposal Disposal Method As discussed above, three different methods are commonly used for spacecraft disposal in Low Earth Orbit. For this reason, a trade-off is conducted to select the most feasible option. Three main parameters are selected for the study: Fuel mass in kilograms considering the mass of the fuel required for end of life manoeuvring. As the mass of the satellites is limited to 150 kg, extra fuel would result in reduced capabilities of the spacecraft as well as issues related to launch procedures. For this reason, Fuel Mass parameter is allocated 80% of the trade-off performance. Re-entry Period in years considering the period required spacecraft disintegration. Longer periods would result in operational complexities and higher risk of debris generation. Population Danger is a factor related to the risk of space debris hitting populated areas. Based on the definition of the parameters, a method scoring lower would be the preferable option for end of life operations. Table 8-1 provides the results obtained from the trade-off performed for constellation disposal. Table 8-1 Constellation disposal trade-off Parameter & Importance Fuel Mass (kg) 80% Re-Entry Period (Years) 10% Population Danger 10% Score (Lower ==> Better) Controlled Re-Entry (3 hours) Graveyard Orbit Uncontrolled Re-Entry Calculation example for uncontrolled re-entry: 65

78 End of Mission Considerations Score = Fuel Mass Importance + Re Entry Period Importance + Population Danger Imporatance = = In terms of fuel mass and danger to Earth s population, it can be seen that disposing the satellites into graveyard orbits is the most feasible option. However, due to the size of the constellation and its inclination of 60 degrees which is generally used by communication and navigation MEO satellites, graveyard disposal is not feasible for the mission design. Based on the results obtained, controlled and uncontrolled re-entry methods score close results. However, the most feasible option is to allow atmospheric drag to decay the satellite orbit as this would provide more flexibility with design and launch operations Computation Following the trade-off performed above, a computation is performed in order to find the orbital parameters ensuring satellite re-entry complying with IDAC requirements. The computation performed is based on that atmospheric drag acts opposite to spacecraft motion and thus reducing its orbital energy. The equation for acceleration due to drag can be defined as: a D = 1 2 ρ C D A m V2 where ρ is the atmospheric density, C D represents the dimensionless drag coefficient of the spacecraft, A stands for the cross-sectional area of the spacecraft, m is the mass of the satellite and V represents the velocity of the spacecraft. International Standard Atmosphere (ISA) defines the rate air density decreases with altitude as follows: δh ρ ρ 0 e h 0 66

79 End of Mission Considerations where ρ and ρ 0 represent atmospheric density at any two altitudes, δh is the difference in altitude between those the altitudes considered and h 0 is the atmosphere scale height, by which air density drops by approximately 1/e. h 0 is usually set to be a value lying in between 50 km and the Karman line (approximately 100 km). Of course, atmospheric height is not constant throughout the globe and throughout the year. Atmospheric height is influenced by many parameters, the most important of which are gravitational fields of the Earth and Solar activity. For uncontrolled orbital re-entries, a gravity standard can be used. Similarly, mean values for the solar activity can be assumed when satellite re-entry is lasting over the Sun cycle of 11 years. Satellite mass is directly related to the re-entry time span of an orbiting body as heavier objects carry more KE. Thus, heavier satellites require additional drag to slow the spacecraft velocity meaning elongation in the decay period. Spacecraft mass, cross sectional area and drag coefficients are the parameters defining the so-called spacecraft ballistic coefficient used for orbital decay computations. Ballistic coefficient of a body re-entering the atmosphere can be defined as follows: C B = m C D A Standard circular orbits in LEO tend to degrade evenly due to the consistent atmospheric drag. However, highly elliptical orbits used for uncontrolled re-entry from higher altitudes tend to degrade differently. Due to that drag acts at perigee and when the spacecraft is in lower altitudes, apogee decay is predominant, while perigee decay is minimal. This sequence defined as aero braking with atmospheric skip is repeated up to the point eccentricity is significantly reduced with apogee in a denser atmosphere when the satellite eventually re-enters. Computational models for uncontrolled re-entry are based on calculations approximating changes in the semi-major axis, a and eccentricity, e. The changes of those orbital parameters due to drag during an orbit can be approximated to: 67

80 End of Mission Considerations δa rev = 2π C D A m a2 ρ p exp( c) [I 0 + 2eI 1 ] δe rev = 2π C D A m a ρ p exp( c) [I 1 + e I 0 + I 2 2 ] where ρ p is the atmospheric density at the perigee of the ellipse, c ae/h, h represent the atmospheric height and I j are the modified Bessel functions of order j and argument c. Based on the method discussed above, computations and simulations were performed in order to find the final orbital parameters ensuring atmospheric reentry within 25 years of mission final activities. The key parameters used for the computation are: Initial apogee altitude 2000 km. Drag coefficient 2.2 (standard for satellites with a cubic body). Cross-sectional area 2.67 m 2 (spacecraft body and solar panels multiplied by a factor of 0.75 taking into account change in attitude due to atmospheric forces). Satellite mass kg. It must be noted that the computation performed considers orbital manoeuvres and fuel burn rates. Initial satellite mass is assumed to be 150 kg. Atmospheric scale height 53 km (lower for safety reasons). Solar Activity mean values. Orbital decay requirement 25 years. Earth s gravitational constant km 3 /s 2. Earth equatorial radius km. The result obtained show that placing a satellite with the above parameters into a 2000 x 429 km orbit would result in natural decay due to drag in about 25 years. To be precise, the orbital parameters obtained show that the spacecraft would be within the region of 30 km in altitude meaning that all spacecraft mass should have burned up by this point. The results obtained comply with IADC-02-68

81 End of Mission Considerations 01 Revision 1 from September 2007 and with ISO 24113:2011 published by the International Organization for Standardization and prepared by Technical Committee ISO/TC 20, Aircraft and space vehicles, Subcommittee SC 14, Space systems and operations. Figure 8-2 provides a graphical representation of the orbital decay showing significant changes to the apogee and minimal changes to the orbital perigee, up to the point both values approach merging point. Observing the figure and the results computed, at 24 years and 9 months apogee and perigee align after which orbital decay is relatively constant for both parameters. Figure 8-2 Orbital decay 69

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83 Attitude Determination and Control System ADCS 9 Attitude Determination and Control System ADCS The purpose of the ADCS systems is to determine and control the attitude of the satellite with high precession. The high pointing accuracy required by the antenna payload makes the ADCS system an important part of the overall design. The current section gives an overview of ADCS proposed design. 9.1 System Overview The designed satellite is a small to medium, nadir pointing, telecommunication satellite orbiting around earth at low earth orbit (LEO) at approximately 2000 km. During the preliminary design review (PDR) the ADCS system requirement were provided by the systems engineering and are summarized in the following table. Table 9-1 ADCS system requirements 9.2 ADCS Modes Description Value unit Pointing Accuracy (Nadir) <0.5 Deg Power budget 25 Watt Mass Budget 20 Kg Detumbling and Data Acquisition Mode After separation from the launcher, a residual torque may exist and needs to be cancelled by detumbling the spacecraft. The ADCS sensors are calibrated and an acquisition of the Sun relative position is performed before using any instrument that could be degraded by an unintended pointing toward the Sun. This mode requires the use of inertial sensors and thrusters Normal Mode This is the default mode of the ADCS. The system must maintain the pointing accuracy of the antennas and compensate any environmental disturbances. Moreover, the satellite will be oriented appropriately to maximize the solar energy absorption by the solar panels. 71

84 Attitude Determination and Control System ADCS Orbit Correction Mode Maintain the spacecraft on the right orbit is necessary and this mode deals with allowing the use of the main propulsion system by pointing in the right direction before any main boost. Moreover, error in manufacturing could lead to a misalignment of thrusters and centre of mass that would create a torque on the spacecraft's body during initial boost and need to be cancelled by the ADCS system Safe Mode The safe mode is enabled in case of failure of a control instrument. The ADCS shall be able to get an attitude to meet the minimal housekeeping requirements in terms of power, communication, safety of instruments and orbit. 9.3 Design Considerations In order to select appropriate control and sensor components, the main factors that constrain the overall design must be quantified. Therefore, the worst-case environmental disturbances were calculated as well as the pointing accuracy requirement. Environmental disturbances In Low-Earth orbit mission, environmental disturbances will affect the satellite s attitude through gravitational and magnetic forces or drag and pressure imposed by external particles. These effects produce torques or velocity changes which must be countered to achieve the right pointing or perform an accurate maneuver. The main environmental disturbances usually considered are listed below and impact the spacecraft's attitude relatively to its location and pointing direction in space: Gravitational effect Magnetic waves Solar radiation Aerodynamic drag 72

85 Attitude Determination and Control System ADCS The ADCS system has been appropriately designed to compensate for the disturbances that the satellite will experience at the reference orbit. The worstcase conditions were considered. 9.4 Hardware Selection The designed satellite is three-axis stabilized satellite. The main trade-off hardware sizing parameters are the ones related to power efficiency, mass and lifetime. The important requirements are the torque efficiency and the pointing accuracy because of the payload requirements of the mission that highly rely on these two performances. The trade-off analysis resulted that regarding the actuators, reaction wheels were selected for their high accuracy supported by thrusters to be able to perform agile manoeuvres and desaturate the wheels when required or in case of emergency. With regards to attitude determination sensors, the high accuracy of star trackers makes them essential for the mission objectives coupled with gyroscopes that provide inertial measurements. For redundancy, safe mode and when star trackers cannot be used (because sensitive to bright stars and inefficient when spinning), sun sensors have been selected. The selected attitude control hardware is presented bellow along with its characteristics Reaction Wheels Smooth changes in torque allows very accurate pointing Nominal speed = 0 Risk of saturation A typical reaction wheels configuration is a pyramid of 4 identical wheels, optimized configuration for redundancy (3-axis control still achievable even in case of failure of one of the wheels), symmetry and balanced torque around any axis. 73

86 Attitude Determination and Control System ADCS Four reaction wheels were selected, of maximum torque provided equal to 0.04 Nm and total angular momentum of 1.5 Nms Thrusters High-torque application No power needed but propellant Provide desaturation of wheels Thrusters are the most frequently used attitude actuators because of their dual use in adjusting orbital parameters such as controlling attitude, nutation, spin rate, performing large and rapid slews and managing angular momentum. An important advantage of using them as reaction controllers is the high level of torque that can be obtained, dearly needed in certain tasks. Nevertheless, attitude control accuracy is directly determined by the minimum thruster impulse available and cannot meet high pointing accuracy requirements for typical communication applications. Usually, at least six thrusters are used to provide control on any axis. For our mission purpose where ADCS thrusters are used only during detumbling mode, safe mode and for reaction wheels desaturation, the use of only four thrusters can achieve the same space manoeuvres. Four thrusters mounted symmetrically about the spacecraft's center of mass provide control torques about all three axes. A tetrahedral configuration of the thrusters allows both attitude control and orbit control to be achieved using the same set of thrusters and makes calculations easier because of equal torque arms along each axis. Four 10N bi-propellant thrusters have been selected and the propellant required for the mission has been estimated to 3.5 Kg plus 20% of margin finally equal to 4.2 Kg. For attitude determination, the following sensors were selected and are presented bellow along with their characteristics. High quality sensors are required to perform accurate pointing in order to limit noise and measurement errors. 74

87 Attitude Determination and Control System ADCS Star Trackers Very high accuracy Adapted to 3-axis stabilized spacecraft The vehicle must be stabilized before operating efficiently Sun Sensors Detectors popular, accurate and reliable Require a clear field of view Very small and inexpensive Usable in low power acquisition and fault recovery modes Avoid other sensors to point toward bright stars Gyroscope Inertial sensor that measures speed and angle changes To be used coupled with external sensors Used for nutation damping and attitude control while ring Smooth and high frequency information (using with external sensor) High accuracy external sensors (star trackers and sun sensors) have been selected and will work coupled with inertial sensors (gyroscopes) for maximum performance. For redundancy, because they are sensitive instruments, 2 star trackers are used and will not operate at the same time. Six sun sensors are used, one located on each face of the spacecraft. Finally, one gyroscope MEMS are used on each axis for a full rate measurement. The following tables provide summarised information of the ADCS hardware properties. 75

88 Attitude Determination and Control System ADCS Table 9-2 ADCS hardware properties part 1 Subsystem Number Dimensions (mm) Mass (Kg) Power (W) Temperature (deg C) Reaction wheels 4 150x150x Thrusters 4 Nozzle d Propellant Star trackers 2 120x120x Baffle 2 d234x Processor 2 245x165x Sun sensors 6 95x107x Gyroscope 3 D365x Table 9-3 ADCS hardware properties part 2 Subsystem Accuracy (deg) Torque available (Nm) Reaction wheels Thrusters 1 10 Star trackers on pitch, yaw on roll Sun sensors 1 - Gyroscope 0.5/sec - 76

89 Electrical Power Subsystem 10 Electrical Power Subsystem This section describes the Electrical Power Subsystem (EPS) of the mission and states the requirements as well as all deign considerations and decisions Power Requirements The EPS has been designed to generate, store, distribute and control power to all the satellites in the constellation. In order to ensure the EPS is designed to a high standard, it has to fulfil the following requirements: The subsystem shall provide solar and battery power to each satellite during the entire mission at all stages. The subsystem shall provide a control method for the power distribution to other satellite subsystems. The subsystem shall provide the satellite with a safe mode that must be able to keep the payload and satellite in a functional state. The subsystem shall provide health and status data of power usage and battery status to the On-Board Data Handling system Power Budget The average and peak power required by one satellite is shown below in Table

90 Electrical Power Subsystem Table 10-1 Power budget Subsystem Average Power Peak Power Average Power + Peak Power + (W) (W) 10% margin (W) 10% margin (W) Payload Structure Thermal Power TT&C OBDH AOCS Propulsion Total The table shows an average operational value of W for the satellite. The peak power shows the maximum power expected to be consumed by each subsystem. The satellite is capable of generating W BOL and W EOL. The EPS provides power to the other subsystems. The Structures and Mechanisms subsystem will need power for the movement of the antennas and solar arrays; battery power will be used at the beginning of the mission to deploy the solar arrays and antennas. Although, the thermal system for the mission is passive, power is needed for the electronically controlled heaters and temperature sensors. Thermal power will vary for sunlit and eclipse periods (where thermal control is more critical). Attitude, Orbit and Control subsystem (AOCS) comprises of reaction wheels, magnetometer, torque rods, star trackers, GPS receiver, inertial measurement unit earth and sun sensors that require electrical power. The Propulsion subsystem will need power for the valve, turbo pumps and any other mechanical devices. The On-Board Data Handling (OBDH) system uses less power for housekeeping than during processing duties. Power allocated to the EPS is for loss of power through the electrical harness, interconnects and power consumption for the power conditioning and distributing 78

91 Electrical Power Subsystem unit. A 10% margin is added to the budget to account for components being modified from a heritage design. The margin also accounts for any power losses that could occur during the transfer of electrical power Power Generation The main source of power is dependent on the operational environment of the satellite and the lifetime of the mission. As the constellation is operating in LEO, this automatically rules out the use of nuclear power, primary batteries and fuel cells. The main source of power therefore will be solar photovoltaic solar arrays Primary Power Gallium Arsenide NeXt Triple Junction Prime Solar Cells from SpectroLab have are most likely to be used as they have a beginning-of-life (BOL) efficiency of 30.7% which significantly reduces the solar array area needed. These cells are also capable of delivering 66,060 cycles and are extremely lightweight with a mass of 2.06 kg/m 2 [18]. A more detailed specification of the intended solar cell can be seen in the Appendix Secondary Power During periods of eclipse, the solar arrays will not be able to provide power to the satellite. To compensate this, batteries will be used instead. Saft VL51ES Lithium- Ion rechargeable cell has been chosen due to its high specific energy of 51 Ah, 79

92 Electrical Power Subsystem very efficient, low weight and of course, the quality of Saft products in general. The VES 16 cell specification can be seen in the Appendix. Figure 10-1 Li-Ion batteries [19] The cells will be placed in a battery module that will contain its own heaters as well as a telemetry interface with the Power Conditioning and Distribution Unit (PCDU) to provide status, battery voltage, current and temperature data Power Distribution, Management and Control Electrical power will be transferred to a regulated 28 V satellite bus using a direct energy transfer system. This system will ensure that extra power is dissipated at BOL and that power is safely transferred. Power will be distributed through an electrical harness that is insulted to keep heat dissipation to a minimum. 80

93 Electrical Power Subsystem Figure 10-2 SST Power conditioning and distribution units [20] Power will be managed with a PCDU from Surrey Satellites as shown above. The unit will be radiation-hardened will protect the system from overcharging, overdischarging, and overheating of the batteries and other satellite subsystems. In the event of failure, the fault detection circuits in the system will be pinpoint where the fault has occurred. Then the EPS will isolate components by cutting off power supply using latching current limiters EPS Mass Budget The EPS is comprised of four main components. In order to keep within the mass budget that has been specified, an estimated mass budget for the power subsystem has been conducted and can be seen in the table below. Table 1.2 Power mass budget Power Mass Budget Units (kg) Solar Arrays 5.02 PCDU 4.40 Harness 3.00 Battery module 8.59 Total EPS mass These estimations have been calculated based on the mission parameters. For example, it has been calculated that approximately 8 battery cells (including 81

94 Electrical Power Subsystem redundant cells) will be needed in order to provide 182 Wh that is needed during the eclipse period. The solar array mass has been estimated from the array area that is needed to provide around 300 W of electrical power. 82

95 On-Board Data Handling Subsystem 11 On-Board Data Handling Subsystem This chapter talks about the design of the On-Board Data Handling (OBDH) system of the mission Requirements The OBDH system for this satellite has to meet the following requirements in order to ensure efficiency and reliability: The subsystem shall provide a method of monitoring satellite subsystems by collecting systems telemetry. The subsystem shall securely store all payload and housekeeping data. The subsystem shall be capable of recovering from single event upsets. The subsystem shall send telemetry from the satellite subsystems and deliver it to the communications system. The subsystem shall receive telecommands from the communications system and deliver it to the satellite subsystems OBDH Design The OBDH system is important as it is responsible for providing command and control of all the satellite subsystems of the satellite platform and it also commands the payload operations. Therefore, the architecture and hardware for this mission have been carefully chosen. The expected mass for the OBDH is 5.7 kg. This mass comprises of the data bus, the on-board computer (OBC) and its modules and wiring and structure Architecture & Hardware The satellite will have serial bus architecture as it is very reliable and is widely used in the space industry. This architecture allows direct interfacing with the OBDH system and the rest of the satellite subsystems. 83

96 On-Board Data Handling Subsystem The OBC will be centred around a LEON3 microprocessor with a MIL-STD-1553B data bus connecting all systems to the OBC. This bus has been selected as it has extensive flight heritage and is capable of delivering high data rates Memory The OBD will have a dual-redundant memory system as well as random access memory (RAM) storage that will be used for error detection and correction (EDAC). The main memory storage will include 16 GB of storage for the payload and a separate memory of up to 1 GB for housekeeping and utilities data Protection and Fault Tolerance The altitude at which the satellites will be operating in are not subjected to intense radiation, but it is still important that the components are well protected. This is done using All the modules and equipment in the OBDH system and the OBC will be radiation-hardened or radiation-tolerant. In order to monitor the system well, the system will periodically perform a cycle of reading, voting and repairing of memory. This is done to prevent a single event upset from changing data. A voting system will be implemented in which an error in one of the three data packets can be corrected through a majority vote. This EDAC method and is done to prevent bit errors occurring on the same data Subsystem Interfacing Communication between the ground and the satellite is very important. The OBDH system allows data to be received and transmitted between the satellite subsystems. Upon uplink, the OBDH system will receives telecommands data from the ground. The commands will be sent directly to the OBC to be decoded. Then it will be sent through the data bus and directly to the desired subsystem interface. Commands from the ground could include codes to control the AOCS components and sensors. The OBDH system will uplink time-tagged and position-tagged commands which will be stored and executed at a specific time 84

97 On-Board Data Handling Subsystem or when the satellite is in a certain position. The most important interface is between the OBDH system and the communications receiver, to allow data to enter the satellite. Figure 11-1 Example of OBDH interface [21] The communications system will downlink payload and telemetry data to the ground. The OBDH system interfaces with the communications transmitter to allow data to leave the satellite. The OBDH system interface with the AOCS is to collect orbital data. Interface with the EPS to gather information from the PCDU and from the Thermal system sensors to get temperature data. 85

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99 Structure and Configuration 12 Structure and Configuration 12.1 Introduction This section will explain the configuration of the structure of the satellites with its main properties and capabilities, and also the main mechanisms that are used for movable and deployable components Subsystem Requirements These are the requirements related to this subsystem that are derived from the top-level requirements of the mission: The mass of the satellites shall not exceed 150 kg. The mass of the structure and mechanisms shall not exceed 20 kg SSTL-150 Satellite Platform In order to ensure a feasible mission concept, looking for existing designs that use proven technologies is a good way of doing it. For the structure of the satellites, since the restriction in terms of mass are quite demanding, an existing satellite platform from Surrey Satellite Technology will serve as starting point [22]: Figure 12-1 SSTL

100 Structure and Configuration The SSTL-150 is a satellite platform of 153 kg designed for LEO missions that offers a mass of 50 kg for payload [23]. These are some of the characteristics of this satellite: Table 12-1 Main characteristics of the SSTL-150 Peak power (EOL) Maximum Payload Mass Bus Dry Mass Mission Design Life Types of orbits available External Payload Volume Internal Payload Volume Structure Nominal schedule from Order 100 W 50 kg Total price $18,315, kg 7 years LEO 400 km to 1000 km, any inclination 730 mm x 455 mm x 774 mm mm x mm x mm Aluminium and aluminium-skinned honeycomb panels 24 months to payload integration, 31 months to launch Apart from those 3 kg that are exceeded, other aspects of the configuration will need to be changed in order to meet the mission requirements, like the mission lifetime, which has to be 8 years instead of 7. Since the satellites of this mission are going to be communications satellites, the mass that is saved for payload will be used for the subsystems most related to the key drivers of the mission, which are the communications subsystem, the data handling, and the power system. Since the required power for this mission is three times the one provided by this platform, the solar panels will definitely need to be greater Structure Introduction The design of the structure is very simple. It has an almost cubic shape made with aluminium-skinned honeycomb panels. 88

101 Structure and Configuration In the mass estimation breakdown, the mass for the subsystem is of 20 kg taking margins into account, so that value has not to be exceeded by both structure and mechanisms. This section will cover every aspect of the structure in terms of design, material properties and impact protection capabilities Configuration The configuration of the structure is very simple as it has been mentioned. The external dimensions are exactly the same as the SSTL-150 [23]. Figure 12-2 Basic CAD model of the structure The interior is divided into two spaces in order to separate subsystems that have to be outside such as the communications and AOCS from others like data handling, which in the case that are not radiation-hardened have to remain inside the structure for thermal and radiation protection. It is possible then to dispense with one of the panels to leave extra space for the antennas or the thrusters. The interior panel has a whole in order to leave space for the fuel tanks to be in the centre of mass of the satellite, so it does not move once the fuel starts to be consumed. 89

102 Structure and Configuration Table 12-2 External dimensions of the structure Dimensions (mm) Table 12-3 Structure breakdown Thickness (mm) 25 Panel Dimensions (mm) Volume (m 3 ) Quantity Mass (kg) Top 650 x Side A 775 x Side B 650 x Base 640 x Total The dry mass of the basic structure is kg using only aluminium-skinned honeycomb panels with a density of 163 kg/m 3 [24]. Table 12-4 Internal volume calculation Available volume (Truncated pyramid) Property Dimensions Area Volume (mm) (m 2 ) (m 3 ) Top 625 x Base 625 x Height Total

103 Structure and Configuration Figure D views with dimensions of the structure The approximated internal volume is 0.98 m 3, almost one cubic meter. The following table shows the available dimensions for the subsystems: Table 12-5 Available internal dimensions Available Dimensions for Subsystems (mm) Available Dimensions for Payload (mm) x 327 x x 327 x

104 Structure and Configuration Material The structure is made by aluminium-skinned honeycomb panels, which core has the following properties [25]: Table 12-6 Aluminium-skinned honeycomb main properties Thickness (Microns) 70 Ø honeycomb (mm) 3.2 Density (kg/m3) 163 Compressive stabilised strength (MPa) 10.2 The panels are supported by Aluminium 6061-T6 bars that, together with the interior panel, add stiffness to the structure Configuration Introduction In order to allow the control systems of the satellites to work properly, it is necessary to integrate the different components of the structure in order to estimate the centre of mass and the inertia matrix of the body in the most precise way possible. Figure 12-4 CAD model of the satellite 92

105 Structure and Configuration This model includes most of the subsystems with their final mass and dimensions values: Propulsion Attitude Determination and Control System Structure and Thermal Protection Power System Communications and Data Handling are not included for the inertia matrix estimation, since the designs for these subsystems were not finished yet. Figure 12-5 CAD model with systems breakdown The representations of the Communications infrastructure consist in four parabolic reflectors and one phased array but the designs are not accurate. The Figure 12-5 highlights the most important parts for the inertia matrix. The results that CATIA V5 provides with this model is shown in the next table: Table 12-7 Total mass and surface implemented in CATIA V5 Mass (kg) Surface (m 2 )

106 Structure and Configuration Table 12-8 Inertia matrix calculated by CATIA V5 IXX (kgm 2 ) 8.82 IXY (kgm 2 ) IXZ (kgm 2 ) 1.41 IYX (kgm 2 ) IYY (kgm 2 ) IYZ (kgm 2 ) 0.33 IZX (kgm 2 ) 1.41 IZY (kgm 2 ) 0.33 IZZ (kgm 2 ) Figure 12-6 View of the CAD model of the satellite As the Table 12-7 shows, the mass taken into account in the CAD model for estimating the inertia matrix is only 130 kg, so 20 kg are missing from the total mass budget. Those kilograms correspond to the Communications and On-Board Data Handling mainly, but also to the rest of subsystems, since the implementation has been done in a simplified way, but ensuring enough accuracy to design a control system capable of working properly Mechanisms Introduction In order to allow the application of external infrastructure like the solar panels at the same time that is achieved the maximum number of satellites launched per 94

107 Structure and Configuration vehicle, deployment mechanisms are necessary, so the satellites occupy the minimum volume possible inside the fairing Solar Array Deployment Mechanism Due to the dimensions of the solar array (0.68 x 4.21 m), it is critical that they are correctly stored during launch and deployed once the satellite is in the orbit. The system applied will be the Fold-and-Roll-up Blanket with a Deployable Boom. Defined by a study of the Defence Evaluation Research Agency and the University of Cambridge, this system adapts properly to this case [26]. This mechanisms relays on a boom and the capability of the panels of being folded: Figure 12-7 Double fold and roll-up solar array. Image from the University of Cambridge. The solar array is folded over twice with the end bars, and then it is rolled over a roller. A tubular boom serves as a deployable backbone. The proposed type by the document from the University of Cambridge is a Rolatube composite boom in order to reduce the size and mass of the deployment cassette [26]. 95

108 Structure and Configuration 96

109 Thermal Control Subsystem 13 Thermal Control Subsystem 13.1 Mission Drivers for Thermal Design Overall Mission Requirements Being able to connect anyone, from anywhere on Earth has become necessary to ensure development of services and countries all around the world. This is why global coverage is a major concern since many parts of the world cannot currently access suitable data rate in order to exchange fast enough. A constellation of telecommunication satellites can, located on the right orbits (cf. Constellation WP), provide a worldwide global coverage of 50Mbp. This the main goal of the CRANSED Team: designing a constellation of satellites able to meet the following main requirements: Table 13-1 Functional requirements of the thermal control subsystem Functional requirements Shall be capable of delivering 50 Mbps of data connectivity Shall provide continuous global coverage Shall provide inter-satellite communications Shall be able to maintain their orbital station Shall be able to close the communication link to small antennas in the ground TRL shall be not less than 7. The satellite shall be flexible enough to cope with different customer needs. Satellite lifetime will be 8 years based on cost estimations. Satellite s weight shall not exceed 150 kilograms Satellites shall not interfere with others in the GEO Constellation and global coverage shall be available by Thermal Requirements As a low-earth constellation mission, the CRANSED satellites undergo harsh thermal conditions (flux from Earth and Sun) and a thermal design is required to ensure functional use of payload and subsystems. Thermal Control Subsystem (TCS) must maintain the temperature of the components of the spacecraft, and a temperature range in it. Its design highly 97

110 Thermal Control Subsystem depends on the type of mission and hence, it is important to have an accurate idea of planned orbits and components characteristics to achieve a good design. Therefore, it is necessary to know the environment in which the spacecraft will operate and the temperature requirements for the different components to design a functional subsystem. The constraints for the Thermal Control Subsystem are the ones derived from the other subsystems, i.e. temperature requirements. Range of temperature can be very different regarding the components or if it is located internally or externally. Most of the components located outside the spacecraft have a wider range of temperature limitations and can endure tougher space conditions. Most sensitive components to temperature variations can be located in positions on the spacecraft that provide thermal stability. The following table summarizes thermal constraints related to each subsystem: Table 13-2 Temperature requirements for each subsystem Subsystem Minimal Temperature ( C) Maximal Temperature ( C) Internal location Antennas Solar arrays External location AOCS PCDU Batteries 0 50 MMH Fuel N2O To estimate the thermal characteristics, the spacecraft will be sized under the worst cases and hence we need to analyse these cases. First, the thermal characteristics of every space environment, that the spacecraft will encounter, need to be analysed. These different environments depend of: 98

111 Thermal Control Subsystem The power dissipated inside the spacecraft. This depends on the mission phase and the instruments that are in use at this moment. The distance to the Sun that will be constant in our case and equal to 1 AU. The orbited body and hence the amount of albedo and IR flux received from them. In a constellation case, only the Earth is considered. The distance to the orbited body. Again, the further the spacecraft orbits from the orbited body the less albedo and IR flux influence. The orbit has been designed to be located at 1000 Km from the Earth surface. Critical mission phases must be detailed to obtain the worst hot case and the worst cold case. The most critical phases, related to thermal, are the ones which present a high amount of power dissipated (hot case) or a low amount (cold case) and they are executed near to the Sun (hot case) or during eclipse (cold case). Thus, the hottest case will be the one with the maximum input heat and the coldest one will be the one with the minimum Thermal Modelling This section presents the steps taken to achieve the final design of the TCS. First, a preliminary design will be determined; the worst cases of the mission will be established along with a selection of the coating material. Once all the requirements are known, it will be obtained the required temperatures by balancing the heat dissipated inside the spacecraft through the radiator against the heat absorbed by it. There are three kinds of external fluxes: Solar flux, infrared flux and albedo. These last two are only considered when the spacecraft is flying close to a body like in our constellation case. 99

112 Thermal Control Subsystem Figure 13-1 Fluxes impacting LEO satellite The balanced thermal equation used can be written as: Where: Ja, Js and Jp are the albedo, the Solar flux and the IR flux, respectively. Aalbedo, Aplanetary and Asolar correspond to the areas which receives different fluxes. Asurface is the area which radiate heat. a and e are the optical properties of the coating materials, absorptivity and emissivity. s is the Stefan-Boltzmann constant: 5.67x10-8 W/m2K4 Q: is the total heat dissipated inside the spacecraft. The Table 13-3 presents a summary with the critical phases, their distance to the Sun and the power dissipated. 100

113 Thermal Control Subsystem Table 13-3 Summary of critical phases, distances and power dissipated Mission Phase Worst Hot case: Sunlight + High payload use Worst cold case: Eclipse + Low payload use Solar flux Earth Albedo Earth IR Power Dissipated (W) Typical material parameters considered for calculation: Table 13-4 Material characteristics Coating Material Absorptivity Emissivity 2mil Aluminized Teflon mil Quartz Mirrors mil Aluminized Kapton mil Aluminized Kapton Honeycomb aluminium panel Table 13-5 Physical characteristics view factors F (earth face) F (side faces)

114 Thermal Control Subsystem Figure 13-2 Physical characteristics On-board temperatures are obtained using the balanced thermal equation. The following table shows the average temperature obtained for each coating material, in the hot (red) and cold (green) case scenarios. 102

115 Thermal Control Subsystem Table 13-6 Estimated average temperature Solar flux Js Albedo flux Ja Earth IR flux Je Internal heat T( K) Coat material 2mil Aluminized Teflon mil Quartz Mirrors 1 2 mil Aluminized Kapton mil Aluminized Kapton Honeycomb aluminium panel mil Aluminized Teflon mil Quartz Mirrors 1 2 mil Aluminized Kapton mil Aluminized Kapton Honeycomb aluminium panel For structural considerations, honeycomb aluminium panels will be used because of high stiffness and reliability. Finally, we obtain in C: Table 13-7 Satellite temperatures per mission phase Mission Phase Worst Hot case: Sunlight + High payload use Worst cold case: Eclipse + Low payload use Temperature ( C) We can see that for typical environment, none of these coating material is able to keep the subsystems in the right thermal conditions. Indeed, the temperature is always too low, even under sunlight especially with honeycomb aluminium panel 103

116 Thermal Control Subsystem with very low absorptivity. This is why, it is necessary to add systems to heat up the overall spacecraft temperature. Some examples of active control are provided: Heaters: these components are used for heat up different equipment during the coldest case along the mission and therefore avoiding huge fluctuations in temperature. They may include some thermostat to control the temperature of a specific component. Figure 13-3 Heater system Louvers: these active thermal control components allow the rejection of internal heat when they are open and they avoid it when they are closed. So, they can control how much internal heat it is going to be dissipated. Figure 13-4 Louver system Patch heaters will be used for their simplicity (no mechanical actuators needed) and reliability. They will use heating resistors to increase the temperature of the spacecraft. MLI (Multi-Layer Insulation) will be used to limit temperature losses as well as black paint to increase absorptivity and decrease emissivity to have a relatively good efficiency and some thermal inertia while heating up. 104

117 Thermal Control Subsystem Figure 13-5 Multi-layer insulation 13.3 Conclusion Thermal Design To conclude, a simple preliminary thermal design has been done using the thermal balanced equation and the subsystem and honeycomb aluminium is required for structure purposes. The overall temperature of the spacecraft is too low and does not fulfil the payload requirements without the use of active thermal systems even with MLI to limit thermal fluxes with the environment and black paint to slightly change the spacecraft s thermal. This is why patch heaters will be used and consume power to heat up the overall satellite. Typical mass and power budget for simple thermal subsystem: Table 13-8 Thermal control subsystem power and mass budgets Further Development Component Mass (Kg) Power (W) MLI 3 0 Heaters 1 20 paint 1 0 Thermal simulations on software to improve the design as well as testing in vacuum and thermal cycle chamber must be done before flight. 105

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119 REFERENCES REFERENCES 1. Iridium Communications Inc. Iridium Available at: (Accessed: 20 July 2001) 2. Globalstar. Globalstar Available at: (Accessed: 1 January 2017) 3. O3b Networks. O3b Networks Available at: (Accessed: 1 January 2017) 4. OneWeb. OneWeb Wertz JR., Everett DF., Puschell JJ. Space mission engineering : the new SMAD Jin JIN., Linling K., Jian YAN., Xi C., Zuyao NI., Xiaochuan YOU., et al. SSC15-V-2 Smart Communication Satellite ( SCS ) Project Overview. 2014; 7. MIT. Communications Satellite Constellations. Wireless World. 2003; 8. Ericsson. Mobility Report: on the Pulse of the Networked Society. Automatic Documentation and Mathematical Linguistics. 2016; Available at: DOI: /S Perko K et al. X-Band phased array antenna validation report. 2002; 10. Maral, G., Bousquet, M., and Sun Z. Satellite Communication Systems. 5th edn. Chichester: Wiley (ed.) SlidePlayer. Satellite Communication System Available at: (Accessed: 1 June 2017) 12. Yang, Qing, Yu, Wei, Challal Y. A Quasi-Dynamic Inter-Satellite Link Reassignment Method for LEO Satellite Networks. 11th International Conference, WASA. 2016; 13. Tsai M. Link Budget. Available at: 107

120 REFERENCES et.pdf (Accessed: 1 January 2017) 14. Hoffman EJ. Space Communications. Oxford University. 15. International Telecommunication Union (ITU). Article 5 of the Radio Regulations. 2004; 16. International Telecommunication Union (ITU). Regulation of global broadband satellite communication. 2012; 17. National Frequency Planning Group. United Kingdom Frequency Allocation Table. 2013; 18. Spectrolab. Space products: Cells Available at: Saft. MP & SMALL VL Available at: Surrey Satellite Technology US. Power Conditioning and Distribution Unit Datasheet Available at: Addaim. A, Kherras. A and ZEB. Design of Low-cost Telecommunications CubeSat-class Spacecraft. INTECH Open Access Publisher. 2010; 22. NASA. SSTL-150 Satellite Platform. 23. Surrey Satellite Technology. NASA Rapid III SSTL-150 ICD File. 2011; 24. Swiss Composite. Aluminium Honeycomb. 2016; : Rao KK., Rao JK., Gupta KSA. Heat Insulation Analysis of an Aluminium Honeycomb Sandwich. Journal of Thermal Engineering. 2015; 1(3):

121 REFERENCES 26. Pellegrino S., Kukathasan S., Tibert G., Watt A. Small Satellite Deployment Mechanisms. 2000; 109

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123 APPENDICESAtlas V 500 Series Launch System APPENDICES Appendix A Atlas V 500 Series Launch System Figure A 1 Atlas V 500 series launch system 111

124 APPENDICESSolar Cell Datasheet Appendix B Solar Cell Datasheet Figure A 2 SpectroLab 30.7% NeXt triple junction (XTJ) prime solar cells datasheet [18] 112

125 APPENDICESBattery Datasheet Appendix C Battery Datasheet Figure A 3 Saft VL51ES Li-Ion cell datasheet [19] 113