AN INGENIOUS SPACECRAFT ARCHITECTURE FOR INNOVATIVE LOW-COST RESEARCH MISIONS N.

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1 AN INGENIOUS SPACECRAFT ARCHITECTURE FOR INNOVATIVE LOW-COST RESEARCH MISIONS N. Larsson (1,2), R. Lilja (2), M. Örth (3), J. Viketoft (3), J. Bäckström (3), J. Köhler (4), R. Lindberg (4) (1) Tel: (2) OHB Sweden, Viderögatan 6, SE Stockholm, Sweden (3) ÅAC Microtec, Dag Hammarskjölds väg 48, SE , Uppsala, Sweden (4) Swedish National Space Board, Box 4006, SE Solna, Sweden ABSTRACT The Swedish National Space Board has challenged OHB Sweden and ÅAC Microtec to analyse the needs from the Swedish space science community and to define a new microsat architecture for a recurring program of low cost research missions. The result is an ingenious microsat architecture optimized for a range of scientific research missions such as atmospheric, astro and plasma physics. A piggyback launch into Low Earth Orbit is assumed as a baseline considering the total mission cost target and the need to de-orbit the satellite after mission completion. Mission flexibility is provided as the physical satellite architecture is split into three main modules: The payload, the service module and the solar array. The size of the service module is kept to a minimum thanks to COTS electronics and a consistent single string approach. This allows the payload and solar array to utilize most of the allowable launcher volume. Smart reconfiguration of these three modules will allow both earth-facing and space-facing payloads. To meet the demanding pointing and orbit determination requirements, Star Tracker and GPS equipment is included in the baseline design. Potential upgrades of the architecture include formation flying technology and increased downlink data capacity. This paper will present the proposed science missions, the utility based system definition process and the resulting system architecture and development approach. 1 INTRODUCTION Two years ago, at the 4S Symposium in Portorož, the Swedish National Space Board presented a call for ideas to investigate low-cost research missions using small satellites. Twelve full proposals were received and reviewed by the SNSB Space Research Advisory Group (SRAC). OHB Sweden and ÅAC Microtec has during 2013 analysed the needs of all twelve missions and has proposed a new microsat architecture that can be utilized for a program of Swedish research satellites at very low recurring mission cost. Based on the preliminary design of the generic satellite architecture and on the feedback from the industry team, SNSB requested a detailed satellite definition phase to be performed taking into account needs of the five LEO missions originally proposed. The result of this detailed satellite definition study is the InnoSat satellite architecture. 2 SYSTEM ARCHITECTURE AND SYSTEM REQUIREMENTS The InnoSat satellite architecture is designed to meet the requirements for several Swedish scientific research missions. This architecture is defined as all the parts and services needed to fully The 4S Symposium 2014 N. Larsson 1

2 implement an InnoSat science mission, including the spacecraft platform, launch services and operations as depicted below. SPACE SEGMENT Spacecraft GROUND SEGMENT External Ground Station OPERATIONS AND USER SEGMENTS Solar Panels Battery InnoSat Spacecraft Bus MPDU Main Computer 3 Mbps Large S- band Ground Station TM/TC Passage Planning InnoSat Operations Structure and Harness ACS TT&C InnoSat Mission Control Centre LEOP and Commissioning Team Flight SW Misson Control System Routine Supervision Science Payload InnoSat Automation Server User Segment Science Team LAUNCH SEGMENT Launch Vehicle Launch Figure 2-1: InnoSat Top-Level System Architecture The InnoSat team has, as depicted in Figure 2-2, analysed all the proposed 12 missions and concluded that only the three single-satellite Low-Earth Orbit missions in dawn/dusk orbit are possible to implement considering the very challenging cost target set by SNSB. A common set of System Requirements have been established and iterated between the science and industry teams. The main cost drivers have been identified and the science teams have made efforts to relax the satellite requirements in these areas. The industry team has in a similar way adapted the baseline satellite configuration for maximum scientific return within the given cost constraints. Flexibility and scalability potential is provided in the baseline design, so that even more complex and advanced missions can be supported in the future. Figure 2-2: The road to the InnoSat System Specification The 4S Symposium 2014 N. Larsson 2

3 Multi-satellite missions or missions requiring other local solar times will be possible using identical core platform components but with a different satellite physical configuration and additional equipment. The following missions will be possible to implement using the InnoSat platform in its standard configuration: SPHINX: X-ray studies MATS: Mesospheric wave studies The following mission will be possible using the InnoSat platform with an extension for higher electrical power capacity and additional TT&C passages for science data downlink: SIW: Stratospheric wind studies The following mission will be possible using the InnoSat platform core avionics but with a different physical configuration, additional multi-satellite equipment and a non-magnetic attitude control system: Alfvén: Plasma physics 3 SYSTEM DESIGN APPROACH The following system design approach has been employed for the InnoSat re-usable architecture: Small Size: The satellite must be small enough to fit into a piggy-back launch envelope and to allow for low-cost environmental test chambers. Single String: To reduce flight hardware cost and to minimize the software complexity (and thereby minimizing the validation effort), a single string approach is deemed necessary. Pre-defined Attitude Mode: To reduce system complexity for lowest hardware cost, a predefined nominal attitude mode shall be assumed. The most versatile mode, covering all the four considered science missions, is the earth/forward-looking mode. Autonomy: To lower operations costs, a high degree of autonomy shall be implemented (even if this could mean higher NRE and flight hardware costs). Extendibility: It shall be possible to extend the baseline design to cover extended missions (with additional NRE costs): o Sun-Pointing Attitude Mode: For missions with high power demands and/or for astro-physics missions (e.g. stellar imaging) o Agile Attitude Mode: As the sun-pointing configuration but with high re-orientation capabilities. In this configuration, the satellite can be slewed to observe specific targets. o Equipment Redundancy: For longer missions and/or for missions where lower risk is desired. o High Power: Scalability in the power subsystem will allow for more power demanding payloads. o Multi-Satellite: Complementing the architecture with deployment mechanisms for secondary satellites and inter-satellite link capability. The 4S Symposium 2014 N. Larsson 3

4 This system architecture will re-used for each science mission and for a mission meeting the system baseline requirements, only tuning of flight software and re-run of mission-specific analyses will be needed as depicted in Figure 3-1. InnoSat Core Spacecraft Components InnoSat Equipment Baseline OBC SpW Router Mass Memory Equipment Handling SW Transciever S-band Antenna Reaction Wheel Magnetic Torquers GPS Solar Cells TT&C RTU AOCS RTU MPDU TM/TC & OPS SW Sun Sensor Star Tracker Magnetometer Battery Cells Thermal HW InnoSat Reference Designs Structural / Thermal Design Power S/S Sizing Structure Solar Panels InnoSat Baseline Architecture InnoSat Standard Bus InnoSat Mission Control Centre Mission Control System Flight Dynamics System Operational Simulator Automation Server InnoSat Ground Station Electrical Architecture Application Software Reference Payload Harness OBSW Payload Mission Tuning and Spacecraft AIT OBSW Tuning AIT Mech. / Thermal Tuning Ground Segment Tuning Simulator Tuning TM/TC Database Update Misson-Specific Engineering and AIT Complete Mission Specific System Mission Specific Spacecraft Launch Service Flight Operations Misison Specific MCC InnoSat Ground Station or External Service Figure 3-1: InnoSat System Design Approach 4 SPACECRAFT DESIGN 4.1 Key Performance Factors The key performance factors of the InnoSat spacecraft bus is summarized in the table below. Table 4-1: InnoSat Key Performance Factors Satellite mass Size Max payload mass <40 kg 70x65x85 cm Up to 15 kg Max payload power 40 W (orbit average, 06:00/18:00 LTAN SSO ) Design lifetime Downlink bitrate Pointing performance requirements 2 years 3-5 Mbps Max 0.1 deg absolute pointing error Max 0.01 deg pointing knowledge error (reconstructed) Orbit determination Nominal attitude mode On-board GPS Nadir/Forward-looking The 4S Symposium 2014 N. Larsson 4

5 4.2 Spacecraft Physical Design The InnoSat spacecraft bus, depicted in Figure 4-1, has been designed to utilize the most of the available launcher volume available for a piggyback launch, to be optimized for a dawn/dusk sunsynchronous orbit and to provide maximum possibly accommodation volume for the science payload. A single solar panel will provide adequate solar power generation for most missions. For missions demanding even more power, it will be possible to extend the standard satellite bus with two deployable solar panels as shown in Figure 4-2. PAYLOAD VOLUME Figure 4-1: InnoSat standard configuration Figure 4-2: Possible high power extension (with two deployable panels) Depending on the payload being an astrophysics or Earth observing instrument, the payload can be oriented towards zenith or nadir as depicted in Figure 4-3. MATS SPHINX SIW Figure 4-3: Typical payload accommodation (with preliminary inputs from science teams) The 4S Symposium 2014 N. Larsson 5

6 4.3 Spacecraft Functional Design The functional architecture of the InnoSat spacecraft is divided into the classical subsystems as illustrated in Figure 4-4 below. PLATFORM THERMAL CONTROL SUB- SYSTEM (TCS) Heaters Thermistors ELECTRICAL POWER SUBSYSTEM (EPS) Battery S/A Strings Main Power Distribution Unit (MPDU) ATTITUDE CONTROL SUBSYSTEM (ACS) 3x Magnetorquer Power Supervisor DC/DC Converter Low-voltage Power Power Protection and Distribution Electronics Magnetometer DATA HANDLING SUBSYSTEM (DHS) Spacecraft Controller Module (SCM) S/C Controller (urtu) Mission Controller Module (MCM) Distributed Power Control Unit (DPCU) Low-voltage Power and SpaceWire Router Star Tracker GPS 3x Reaction Wheel ACS Interface Electronics Antenna 1 TT&C SUBSYSTEM (TT&C) Switch / Hybride Diplexer Antenna 2 S-Band Transceiver RX TX TTC Controller Module (TCM) TTC Controller (urtu) Mass Memory Unit TTC Interface Electronics DC/DC Converter Science Instrument(s) Payload Controller PAYLOAD Figure 4-4: InnoSat Spacecraft Functional Architecture The Electrical Power Subsystem (EPS) is responsible for power conditioning and distribution of 28 V unregulated power to on-board users and regulated 5 V power to the DHS. The Thermal Control Subsystem (TCS) is responsible for controlling the on-board temperatures and is designed with mainly passive methods and only a limited set of active heaters. The Data Handling Subsystem (DHS) is responsible for telecommand and telemetry handling, for commanding and monitoring of the on-board avionics as well as for executing the on-board software. The Attitude Control Subsystem (ACS) is responsible for controlling the attitude of the satellite using dedicated sensors and actuators. The Telemetry, Tracking and Telecommand (TT&C) Subsystem is responsible for communication with Ground using S-band RF communication. The Payload is the mission-specific instrumentation. The 4S Symposium 2014 N. Larsson 6

7 4.4 Key Design Items Distributed Plug and Play Data Handling System The DHS hardware modules utilize standard components of ÅAC Microtec which on the physical level is interconnected through SpaceWire, Figure 4-5 shows the current µrtu which e.g. the Spacecraft Controller Module (SCM) is based on. On the software level, the data handling subsystem relies on the concepts of the Space Plug-and-Play Avionics (SPA) standard, and more specifically the protocol variant for SpaceWire called SPA-S. This allows for a flexible interconnection of the modules in the DHS and other connected units, while also allowing for a well-defined interface between them and the ability to seamlessly integrate different processing tasks running on the various controller modules. Figure 4-5 Current ÅAC Microtec µrtu FM product Between missions, there is a possibility to leave complete end nodes unaltered and have the mission specific software changes only in one node. This minimizes the changes between missions and reduces cost. The DHS architecture concept of having multiple distributed controller modules gives an advantage that large and vital parts of the on-board software can be left with minimal changes from mission to mission. For an instance, typically the software of the Mission Controller Module (MCM) will be subjected to major changes, whereas the software of the SCM just will have minor changes if an equivalent ACS suite is used. By having minimal changes in the SCM software it can be asserted that core functions such as e.g. de-tumbling, safeguarding and FDIR functions will operate out of the box. With this architecture it is also possible to add and remove controller modules according to mission specific needs. For an instance, if a future mission needs a propulsion system a controller module for such system is connected to the SpaceWire router, this with minimal impact on the rest of the DHS. Two-Mode only Attitude Control System The InnoSat ACS design relies on only two operational modes: The Nominal Pointing (NP) mode is where the scientific observations are made and is solely based on star tracker and reaction wheels. The specific NP mode behaviour for a specific mission is tailored and shaped by mission specific Shaping commands. This is consistent with the DHS architecture and its software, the SCM is to provide an interface for such commands, and the commands themselves will be originated from the MCM. The Intermediate Acquisition and Handover (IAHO) mode provides the Safeguarding and the Handover capacities and is based solely on magnetometer and magnetorquers. The 4S Symposium 2014 N. Larsson 7

8 Power Up OBC reset Reboot cmd INIT Auto Intermediate Acquisition & HandOver (IAHO mode) Auto (Enable/Disable) CMD Auto CMD Ground Station (Operations) Shaping CMDs Nominal Pointing Mode (NP mode) Figure 4-6: The operational modes of the Attitude Control System The IAHO mode is a pseudo three-axis, magnetorquer-only based, attitude control mode. The only sensor used is one (1) 3-axis magnetometer. The control law is based on a minus Bdot type principle complemented with a novel sinusoidal dipole offset algorithm to ensure good illumination of the solar panels even in the absence of a sun sensor. In addition, the IAHO algorithm package can provide a coarse orbit position estimate based on magnetic field analysis. The IAHO mode can provide excellent entrance conditions to the nominal fine-pointing mode. The predicted power profile for IAHO detumbling (5 degrees per second tip-off rate) and establishing a safe state of the S/C is illustrated below. A seen in the graph, the deepest battery DoD during de-tumble is roughly 15% which is well above the acceptable limit. Figure 4-7: Detumbling and safe mode establishment 9 solar strings (the jagged look of the SoC graph is an artefact of the high-fidelity simulator). The 4S Symposium 2014 N. Larsson 8

9 5 DEVELOPMENT AND AIT APPROACH 5.1 Model philosophy The model philosophy adopted for InnoSat is a proto-flight approach where PFM units are used both in the Avionic Test Bench (ATB) and the PFM satellite. The only Engineering Model (EM) units foreseen are those that are identified to have the need of additional testing, typically shock testing. As it stands today only one or two EM units - or more correctly Engineering Qualification Models (EQMs) - will be procured. By undertaking this approach a significant direct cost is reduced in the budget. 5.2 Assembly and integration The concept of a flat satellite has been adopted. This concept gives InnoSat simplicity in both its mechanical design and procedures during assembly and integration. Another design driver in InnoSat is the distinct separation of the payload and satellite platform, this separation provides minimal dependencies between the two, both in terms of complex integration procedures and project scheduling. Another strongpoint of the approach taken is that a design change in the satellite platform or the Service Module as it is called has low likelihood to affect the payload, and wise versa. 5.3 Testing The PFM satellite will be used for environmental testing as well as for final end-to-end validation of the spacecraft functional chains. The flight software and the DHS electronics will mainly be validated using the ATB and the System Simulator. The System Simulator is built from the onboard application software, from equipment simulation models and from environmental simulation models as depicted below. Simulator Simulator AOC Core Application AOC Core Code Equipment SCM Application SW S/C Applications: - System FDIR - Mode Handling - Time Handling - TC Sheduling TM/TC SPA / Msg MCM Application SW Mission Applications: - P/L Handling - P/L Pointing / Shaping cmds Equipment TM/TC SPA / Msg TCM Application SW TM/TC Applications: - TM/TC Routing - MM Handling TM/TC Equipment SPA / Msg AOCS Equipment Models Power / Thermal Equipment Models Simulator Framework TT&C Equipment Models (TBD) AOCS Dynamics & Environmental Models Power Model RAMSES Mission Control System Figure 5-1: System Simulator Architecture On the PFM satellite, the focus will be validation of each functional chain as well as on longduration tests. The 4S Symposium 2014 N. Larsson 9

10 In terms of external test facilities the use of low-cost Swedish test facilities is foreseen. The satellites modest size allows for easy packing and transport between facilities. If cleanliness cannot be guaranteed by the facility, the spacecraft and the payload will be protected by a mobile clean room tent or just encapsulated in double bags, whichever that is the most suitable for the test undertaken. 6 OPERATIONS The objective for the InnoSat operational concept is to minimize costs for the routine management of the spacecraft while the provision of science data is guaranteed on a daily basis. Use of a small, dedicated ground station has been considered but opted out if favour of a procured data service from an external provider such as SSC or K-SAT. The on-board S-band transmitter can then be utilized to its maximum with downlink bitrates up to 5 Mbps thanks to a relatively large ground station antenna. In the routine phase, only one or two passages per day will be sufficient for most science missions. These routine passages will be fully automated in the Mission Control Centre and the ground station provider will be allowed to re-schedule the antenna passages given that the daily downlink data volume is maintained. The backbone of the Mission Control Centre is the RAMSES Mission Control System (MCS), developed by OHB Sweden for the PRISMA formation flying mission. As the MCS is mainly intended for manned satellite operations, it needs to be complemented with automation services in the routine phase. In this phase, the MCC will nominally only be manned once per week for mission planning and satellite maintenance. During all other passages, science data will automatically be downlinked and made available to the science team as depicted in Figure 6-1. Spacecraft TC Onboard Schedule TC Uplink Mission Control Centre InnoSat Automation Server InnoSat XML interpreter (includes repeat validation) MCC Validation Result Payload Coordination Centre XML Validation (through use of XML Schema) Payload Instrument TC Log InnoSat Data Exchange Server Payload TC Plan (XML) TM Onboard Mass Memory TM Downlink Raw Data (RAC) Post Processing Processed Ancillary Data (Format TBD) Coordinated Payload Plan New Development Adapted from Prisma Reuse of existing process/format Orbit Determination Contact Planning Orbit Information (TLE) Spacecraft Contact Schedule (XML?) Science Community Payload Coordinator Figure 6-1: Payload data distribution chain Most of the processes involved are based on processes and interfaces already developed for PRISMA and will require only minor modifications for InnoSat. By introducing a validation step under the responsibility of the Payload Coordination Center (PCC), greater confidence in the payload plan can be achieved prior to uploading it to the spacecraft and instrument itself. 7 CONCLUSIONS We believe that the ingenious InnoSat spacecraft architecture is a perfect match for a program of small, highly capable and affordable Swedish space research missions. Our task has been to enable The 4S Symposium 2014 N. Larsson 10

11 the Swedish space science community to provide first-class science within a break-through low-cost mission budget. This has been made possible thanks to the following key success factors: System Specification: Provides the performance envelope to the science teams and eliminates high non-recurring costs for each mission Optimized for Science: 3-axis stabilized platform with high power and data downlink capabilities. Large, un-obstructed payload accommodation volume. Affordable Launch Solution: The satellite is designed to fit several rideshare launch options. Maximum Heritage Re-use: Plug and play electronics from ÅAC Microtec, application software and mission control system from OHB Sweden. Microsat COTS FM hardware: Now available thanks to many other low-cost missions Single String: Can be justified for short missions and adequate pre-flight testing System Level Trade-offs: Cost-efficient power system, simplified ACS, multi-mission service module Software Simulator: Cost-efficient satellite validation, perfect for operations planning and trouble shooting Automation: Minimal staffing in routine phase, un-manned ground station passages. The 4S Symposium 2014 N. Larsson 11