Earth Observation Satellite

System Design Description (SyDD) — ISO/IEC/IEEE 15289 — Description | IEEE 29148 §6.5
Generated 2026-03-27 — UHT Journal / universalhex.org

System Decomposition

flowchart TB
  n0["system<br>Earth Observation Satellite"]
  n1["subsystem<br>Optical Payload Subsystem"]
  n2["subsystem<br>Attitude and Orbit Control Subsystem"]
  n3["subsystem<br>Electrical Power Subsystem"]
  n4["subsystem<br>Telemetry Tracking and Command Subsystem"]
  n5["subsystem<br>Onboard Data Handling Subsystem"]
  n6["subsystem<br>Thermal Control Subsystem"]
  n7["system<br>Earth Observation Satellite"]
  n8["subsystem<br>Optical Payload"]
  n9["subsystem<br>AOCS"]
  n10["subsystem<br>Electrical Power"]
  n11["subsystem<br>OBDH"]
  n12["subsystem<br>TTC"]
  n13["subsystem<br>Thermal Control"]
  n14["subsystem<br>Structure & Mechanisms"]
  n15["subsystem<br>Propulsion"]
  n0 -->|contains| n1
  n0 -->|contains| n2
  n0 -->|contains| n3
  n0 -->|contains| n4
  n0 -->|contains| n5
  n0 -->|contains| n6
  n1 -->|Image data| n5
  n5 -->|Downlink data| n4
  n2 -->|Attitude telemetry| n5
  n3 -->|Power status| n5
  n8 -->|image data| n11
  n11 -->|downlink data| n12
  n10 -->|regulated power| n11
  n9 -->|thruster commands| n15
  n13 -->|cryo cooling| n8
  n14 -->|array deployment| n10

Earth Observation Satellite — Decomposition

Decomposition Tree

Earth Observation Satellite DDF75219
- Optical Payload Subsystem D4C51018
  - Telescope Assembly CE851018
  - Focal Plane Array D4D51018
  - Image Processing Unit 50F53008
  - Spectral Filter Mechanism D6951008
  - Calibration System 54F73218
- Attitude and Orbit Control Subsystem 55F73218
  - Star Tracker Assembly D5F73218
  - Reaction Wheel Assembly D6F51018
  - Magnetorquer System 54F53208
  - GNSS Receiver
  - AOCS Flight Software 41F73B18
- Electrical Power Subsystem 54D73218
  - Solar Array Assembly CEC51018
  - Battery Assembly D6D51218
  - Power Control and Distribution Unit D7F71008
  - Solar Array Drive Mechanism DFF71208
  - Power Management Software 41B77B18
- Telemetry Tracking and Command Subsystem 54E57018
  - S-Band Transponder D4F57018
  - X-Band Transmitter D4E55018
  - S-Band Antenna Assembly C6851018
  - X-Band High-Gain Antenna DEC51018
  - Baseband Data Processor 51F57008
- Onboard Data Handling Subsystem 51F77018
  - Onboard Computer D1F77018
  - Mass Memory Unit D4851018
  - SpaceWire Network Router D4E57018
  - Remote Terminal Unit D5F57008
  - Onboard Data Handling Software 41F77B18
- Thermal Control Subsystem 51D71208
  - Pulse-Tube Cryocooler D6D51018
  - Multi-Layer Insulation Blankets C6851018
  - Optical Solar Reflector Radiator Panels C6851008
  - Heat Pipe Network C6D51008
  - Thermostatically Controlled Heater System 55F73008
  - Deployable Cryocooler Radiator CE851018
- Structure and Mechanisms Subsystem CE851018
  - Primary Structure CE851018
  - Solar Array Deployment Mechanism CED51218
  - Launch Vehicle Adapter CE851018
  - Antenna Deployment Boom CE951018
  - Harness Assembly C6851018

Subsystem Requirements (SUB)

Ref	Requirement	V&V	Tags
SUB-SUBSYSTEMREQUIREMENTS-001	The Telescope Assembly SHALL provide a clear aperture diameter of not less than 500 mm with a Korsch three-mirror anastigmat design to achieve the required ground sampling distance.	—	subsystem, optical-payload, session-177
SUB-SUBSYSTEMREQUIREMENTS-002	The Focal Plane Array SHALL provide detector elements covering visible (450-690 nm), near-infrared (760-900 nm), and shortwave infrared (1550-1750 nm) spectral bands with quantum efficiency exceeding 0.6 in each band.	—	subsystem, optical-payload, session-177
SUB-SUBSYSTEMREQUIREMENTS-003	While imaging, the Focal Plane Array SWIR detectors SHALL be maintained at an operating temperature of 150 K or below by the integrated cryocooler to achieve the required signal-to-noise ratio.	—	subsystem, optical-payload, session-177
SUB-SUBSYSTEMREQUIREMENTS-004	The Image Processing Unit SHALL implement CCSDS 122.0 lossless compression achieving a compression ratio of at least 2:1 for all spectral bands without loss of radiometric fidelity.	—	subsystem, optical-payload, session-177
SUB-SUBSYSTEMREQUIREMENTS-005	The Image Processing Unit SHALL process and format payload data at a sustained throughput of at least 1.2 Gbps to prevent data loss during continuous imaging passes.	—	subsystem, optical-payload, session-177
SUB-SUBSYSTEMREQUIREMENTS-006	The Spectral Filter Mechanism SHALL achieve out-of-band rejection of at least 10^-4 for each spectral channel to prevent cross-band contamination of image data.	—	subsystem, optical-payload, session-177
SUB-SUBSYSTEMREQUIREMENTS-007	The Spectral Filter Mechanism SHALL complete band selection transitions in less than 200 milliseconds to support multi-band imaging within a single ground track pass.	—	subsystem, optical-payload, session-177
SUB-SUBSYSTEMREQUIREMENTS-008	The Calibration System SHALL provide absolute radiometric calibration accuracy of 5% or better traceable to SI standards across all spectral bands over the full mission lifetime.	—	subsystem, optical-payload, session-177
SUB-SUBSYSTEMREQUIREMENTS-009	While in imaging mode, the Telescope Assembly SHALL maintain primary-to-secondary mirror alignment within 5 micrometres to preserve modulation transfer function performance across the full field of view.	—	subsystem, optical-payload, session-177
SUB-SUBSYSTEMREQUIREMENTS-010	The Star Tracker Assembly SHALL determine inertial attitude with an accuracy of 3 arcseconds (3-sigma) per axis across the full celestial sphere, excluding zones within 20 degrees of the Sun.	—	subsystem, aocs, session-178
SUB-SUBSYSTEMREQUIREMENTS-011	The Reaction Wheel Assembly SHALL provide a total angular momentum storage capacity of at least 50 Nms per axis in a four-wheel redundant tetrahedral configuration to support agile imaging slew manoeuvres of 30 degrees in under 60 seconds.	—	subsystem, aocs, session-178
SUB-SUBSYSTEMREQUIREMENTS-012	The Magnetorquer System SHALL generate a magnetic dipole moment of at least 30 Am2 per axis to desaturate the Reaction Wheel Assembly from 80% stored momentum to below 20% within one orbit period at any orbital inclination.	—	subsystem, aocs, session-178
SUB-SUBSYSTEMREQUIREMENTS-013	The GNSS Receiver SHALL provide autonomous orbit position determination with accuracy better than 10 metres (3-sigma) and velocity determination better than 0.1 m/s (3-sigma) using dual-frequency GPS/Galileo L1/L5 signals.	—	subsystem, aocs, session-178
SUB-SUBSYSTEMREQUIREMENTS-014	The AOCS Flight Software SHALL implement an extended Kalman filter fusing star tracker, GNSS, and magnetometer measurements to provide filtered attitude knowledge with accuracy better than 0.005 degrees (3-sigma) per axis at a minimum update rate of 10 Hz.	—	subsystem, aocs, session-178
SUB-SUBSYSTEMREQUIREMENTS-015	When a fault is detected by the AOCS Flight Software, the AOCS shall autonomously transition from any operational mode to safe mode within 10 seconds while maintaining Sun-pointing attitude to preserve positive power balance.	—	subsystem, aocs, session-178
SUB-SUBSYSTEMREQUIREMENTS-016	While in imaging mode, the Reaction Wheel Assembly SHALL limit micro-vibration disturbance torque to below 10 mNm amplitude at frequencies above 1 Hz to prevent image smear in the Optical Payload Subsystem.	—	subsystem, aocs, session-178
SUB-SUBSYSTEMREQUIREMENTS-017	The GNSS Receiver SHALL provide a UTC time reference with accuracy better than 100 nanoseconds (1-sigma) to the Onboard Data Handling Subsystem for payload image time-stamping and orbit event synchronisation.	—	subsystem, aocs, session-178
SUB-SUBSYSTEMREQUIREMENTS-018	The Solar Array Assembly SHALL generate a minimum beginning-of-life power of 3200 W under AM0 illumination at normal incidence to ensure 2500 W end-of-life orbital average power after accounting for cell degradation, cosine losses, and eclipse fraction over a 7-year mission.	—	subsystem, eps, session-179
SUB-SUBSYSTEMREQUIREMENTS-019	The Battery Assembly SHALL provide a minimum usable energy storage capacity of 1500 Wh at end-of-life to sustain all essential spacecraft loads during a maximum eclipse duration of 35 minutes without bus voltage excursion below 26 V.	—	subsystem, eps, session-179
SUB-SUBSYSTEMREQUIREMENTS-020	The Power Control and Distribution Unit SHALL regulate the primary power bus to 28 V +/- 2 V under all load conditions from minimum keep-alive to peak imaging power, including load transients of up to 500 W step change.	—	subsystem, eps, session-179
SUB-SUBSYSTEMREQUIREMENTS-021	The Solar Array Drive Mechanism SHALL track the Sun with a pointing accuracy of 1 degree or better about the rotation axis throughout the sunlit portion of each orbit to maintain solar array power output within 2% of the maximum achievable.	—	subsystem, eps, session-179
SUB-SUBSYSTEMREQUIREMENTS-022	When the battery state of charge falls below 30%, the Power Management Software SHALL autonomously execute load shedding of non-essential loads within 500 milliseconds in a priority sequence defined by the ground-configurable load priority table.	—	subsystem, eps, session-179
SUB-SUBSYSTEMREQUIREMENTS-023	The Battery Assembly SHALL sustain a minimum of 60000 charge-discharge cycles at a depth of discharge not exceeding 25% with capacity fade of less than 20% from beginning-of-life rated capacity over the 7-year mission duration.	—	subsystem, eps, session-179
SUB-SUBSYSTEMREQUIREMENTS-024	The Power Control and Distribution Unit SHALL implement maximum power point tracking with an efficiency of 97% or better across the full solar array voltage range from beginning-of-life cold to end-of-life hot operating conditions.	—	subsystem, eps, session-179
SUB-SUBSYSTEMREQUIREMENTS-025	When the spacecraft enters safe mode, the Power Management Software SHALL configure the Electrical Power Subsystem to maintain positive power balance by shedding all payload loads and non-essential platform loads within 2 seconds while preserving Sun-pointing attitude control power.	—	subsystem, eps, session-179
SUB-SUBSYSTEMREQUIREMENTS-026	The Onboard Computer SHALL employ a radiation-hardened processor with a minimum computational throughput of 400 MIPS and 256 MB of EDAC-protected SRAM to support concurrent execution of FDIR, command processing, and telemetry formatting tasks.	—	subsystem, obdh, session-180
SUB-SUBSYSTEMREQUIREMENTS-027	The Mass Memory Unit SHALL provide a minimum of 2 Tbit usable storage capacity using NAND flash with triple-module redundancy error correction achieving a bit error rate below 10^-15 after correction.	—	subsystem, obdh, session-180
SUB-SUBSYSTEMREQUIREMENTS-028	The SpaceWire Network Router SHALL provide at least 8 bidirectional SpaceWire ports operating at a link rate of 200 Mbps per port with deterministic latency not exceeding 5 microseconds per hop.	—	subsystem, obdh, session-180
SUB-SUBSYSTEMREQUIREMENTS-029	The Mass Memory Unit SHALL support simultaneous write from the payload data stream and read for downlink at a sustained aggregate throughput of 1.5 Gbps without data loss or write-stall conditions.	—	subsystem, obdh, session-180
SUB-SUBSYSTEMREQUIREMENTS-030	The Onboard Data Handling Software SHALL implement autonomous fault detection, isolation, and recovery covering at minimum single-event upsets, processor watchdog timeout, memory parity errors, and SpaceWire link failures, with fault detection latency not exceeding 500 milliseconds.	—	subsystem, obdh, session-180
SUB-SUBSYSTEMREQUIREMENTS-031	The Remote Terminal Unit SHALL acquire analog sensor measurements with 16-bit resolution and absolute accuracy of 0.1% of full scale, sampling at least 100 channels at a minimum rate of 1 Hz for housekeeping telemetry collection.	—	subsystem, obdh, session-180
SUB-SUBSYSTEMREQUIREMENTS-032	When a telecommand is received, the Onboard Data Handling Software SHALL validate command authentication, sequence count, and parameter range within 100 milliseconds and reject commands failing any validation check with a rejection telemetry report.	—	subsystem, obdh, session-180
SUB-SUBSYSTEMREQUIREMENTS-033	The SpaceWire Network Router SHALL distribute time-code ticks synchronised to the GNSS-derived UTC reference with a maximum jitter of 1 microsecond to all connected subsystem nodes for correlation of payload and housekeeping timestamps.	—	subsystem, obdh, session-180
SUB-SUBSYSTEMREQUIREMENTS-034	The S-Band Transponder SHALL receive telecommands at data rates up to 64 kbps on the 2025-2110 MHz uplink band and transmit telemetry at data rates up to 512 kbps on the 2200-2290 MHz downlink band, with a minimum receiver sensitivity of -110 dBm.	—	subsystem, ttc, session-182
SUB-SUBSYSTEMREQUIREMENTS-035	The S-Band Transponder SHALL support coherent turnaround mode with a transponder ratio of 240/221 to enable two-way Doppler range-rate measurements with an accuracy better than 0.1 mm/s.	—	subsystem, ttc, session-182
SUB-SUBSYSTEMREQUIREMENTS-036	The X-Band Transmitter SHALL deliver a minimum RF output power of 8 W at the antenna port across the 8.025-8.4 GHz downlink band, supporting selectable data rates of 150, 300, 450, 600, and 800 Mbps with OQPSK modulation.	—	subsystem, ttc, session-182
SUB-SUBSYSTEMREQUIREMENTS-037	The X-Band Transmitter SHALL achieve a power-added efficiency of not less than 25% at the rated 8 W output, with total DC power consumption not exceeding 50 W in transmit mode.	—	subsystem, ttc, session-182
SUB-SUBSYSTEMREQUIREMENTS-038	The Baseband Data Processor SHALL implement CCSDS AOS transfer frame assembly with virtual channel multiplexing supporting at least 8 virtual channels, and SHALL apply concatenated Reed-Solomon (223,255) and convolutional (rate 1/2, K=7) forward error correction coding for X-band downlink.	—	subsystem, ttc, session-182
SUB-SUBSYSTEMREQUIREMENTS-039	The Baseband Data Processor SHALL sustain a continuous throughput of 800 Mbps from the Mass Memory Unit to the X-Band Transmitter without frame loss during ground station passes.	—	subsystem, ttc, session-182
SUB-SUBSYSTEMREQUIREMENTS-040	The X-Band High-Gain Antenna SHALL provide a minimum gain of 34 dBi at 8.2 GHz with a two-axis gimbal mechanism enabling ±90 degree steering in both azimuth and elevation to track ground stations during LEO passes.	—	subsystem, ttc, session-182
SUB-SUBSYSTEMREQUIREMENTS-041	The X-Band High-Gain Antenna gimbal SHALL achieve a pointing accuracy better than 0.3 degrees and a slew rate of at least 2 degrees per second to maintain link margin during overhead passes with angular rates up to 1 degree per second.	—	subsystem, ttc, session-182
SUB-SUBSYSTEMREQUIREMENTS-042	The S-Band Antenna Assembly SHALL provide near-omnidirectional coverage with a minimum gain of -3 dBi over at least 95% of the sphere surrounding the spacecraft to ensure telecommand reception and telemetry transmission in any attitude.	—	subsystem, ttc, session-182
SUB-SUBSYSTEMREQUIREMENTS-043	When the onboard schedule initiates a ground station pass, the Baseband Data Processor SHALL prioritise real-time imaging data replay over deferred housekeeping data, ensuring that the highest-priority virtual channel occupies at least 80% of the available downlink bandwidth.	—	subsystem, ttc, session-182
SUB-SUBSYSTEMREQUIREMENTS-044	While the spacecraft is in safe mode, the S-Band Transponder SHALL remain operational on low-power mode with telemetry transmission at a minimum rate of 4 kbps to support anomaly diagnosis from the ground.	—	subsystem, ttc, session-182
SUB-SUBSYSTEMREQUIREMENTS-045	The S-Band Transponder SHALL implement CCSDS Proximity-1 compatible command authentication to reject unauthorized telecommands with a false-acceptance probability below 10^-16 per command.	—	subsystem, ttc, session-182
SUB-SUBSYSTEMREQUIREMENTS-046	The Pulse-Tube Cryocooler SHALL maintain the SWIR focal plane array cold-finger temperature at 150 K ±2 K throughout the orbital cycle, including during eclipse transient conditions.	—	subsystem, thermal-control, session-183
SUB-SUBSYSTEMREQUIREMENTS-047	The Thermostatically Controlled Heater System SHALL maintain battery pack temperature above 0°C and propellant line temperature above 5°C during eclipse and safe-mode conditions without OBC intervention.	—	subsystem, thermal-control, session-183
SUB-SUBSYSTEMREQUIREMENTS-048	The Optical Solar Reflector Radiator Panels SHALL reject a minimum of 600 W orbit-average heat load while maintaining panel equilibrium temperature below 20°C at end-of-life solar absorptance.	—	subsystem, thermal-control, session-183
SUB-SUBSYSTEMREQUIREMENTS-049	The Heat Pipe Network SHALL limit equipment panel temperature gradients to less than 5°C across any 0.5 m span under maximum internal dissipation conditions.	—	subsystem, thermal-control, session-183
SUB-SUBSYSTEMREQUIREMENTS-050	The Deployable Cryocooler Radiator SHALL reject a minimum of 100 W at a panel equilibrium temperature of -40°C to provide adequate thermal margin for end-of-life cryocooler waste heat dissipation.	—	subsystem, thermal-control, session-183
SUB-SUBSYSTEMREQUIREMENTS-051	The Multi-Layer Insulation Blankets SHALL achieve an effective emittance of less than 0.03 over all covered external surfaces to limit parasitic heat leaks to below 50 W total.	—	subsystem, thermal-control, session-183
SUB-SUBSYSTEMREQUIREMENTS-052	While the spacecraft is in safe mode, the Thermal Control Subsystem SHALL maintain all equipment within survival temperature limits using only survival heater circuits powered from the unregulated bus.	—	subsystem, thermal-control, session-183
SUB-SUBSYSTEMREQUIREMENTS-053	The Pulse-Tube Cryocooler SHALL export vibration forces of less than 100 mN at the compressor drive frequency to prevent degradation of optical payload image quality.	—	subsystem, thermal-control, session-183
SUB-SUBSYSTEMREQUIREMENTS-054	The Primary Structure SHALL maintain positive structural margins under combined quasi-static loads of 8g axial and 3g lateral with a factor of safety of 1.25 on yield and 1.5 on ultimate for all load-bearing members.	—	subsystem, structure-mechanisms, session-184
SUB-SUBSYSTEMREQUIREMENTS-055	The Primary Structure SHALL exhibit a first natural frequency above 25 Hz in the axial direction and above 10 Hz in the lateral direction in the stowed launch configuration.	—	subsystem, structure-mechanisms, session-184
SUB-SUBSYSTEMREQUIREMENTS-056	The Solar Array Deployment Mechanism SHALL deploy both solar array wings to the operational configuration within 120 seconds of receiving the deployment command, with deployment confirmation telemetry provided to the OBDH within 5 seconds of latch engagement.	—	subsystem, structure-mechanisms, session-184
SUB-SUBSYSTEMREQUIREMENTS-057	The Launch Vehicle Adapter SHALL achieve clean separation with tip-off rates below 1.0 deg/s in all axes and a relative separation velocity of 0.5 m/s plus or minus 0.1 m/s.	—	subsystem, structure-mechanisms, session-184
SUB-SUBSYSTEMREQUIREMENTS-058	The Antenna Deployment Boom SHALL exhibit a first deployed natural frequency above 5 Hz with the antenna payload mass of 3.2 kg at the tip, to avoid coupling with the AOCS control bandwidth.	—	subsystem, structure-mechanisms, session-184
SUB-SUBSYSTEMREQUIREMENTS-059	The Harness Assembly SHALL maintain a minimum separation of 50 mm between power, signal, and pyrotechnic harness routing paths and SHALL employ shielded connectors with EMI backshells on all data interfaces to limit conducted emissions below MIL-STD-461G CE102 limits.	—	subsystem, structure-mechanisms, session-184
SUB-SUBSYSTEMREQUIREMENTS-060	The Solar Array Deployment Mechanism hold-down and release system SHALL achieve a release reliability of at least 0.9999 per wing using redundant Frangibolt actuators with independent initiation circuits.	—	subsystem, structure-mechanisms, session-184
SUB-SUBSYSTEMREQUIREMENTS-061	The Thruster Assembly SHALL deliver a thrust of 0.9 N to 1.1 N per thruster over the blowdown pressure range of 22 bar to 5.5 bar with a specific impulse of at least 215 seconds.	—	subsystem, propulsion, session-184
SUB-SUBSYSTEMREQUIREMENTS-062	The Thruster Assembly SHALL provide a minimum impulse bit of 50 mN-s or less to support fine orbit adjustment manoeuvres.	—	subsystem, propulsion, session-184
SUB-SUBSYSTEMREQUIREMENTS-063	The Propellant Tank SHALL store a minimum of 45 kg of hydrazine propellant with a MEOP of 24 bar and an expulsion efficiency of at least 98%.	—	subsystem, propulsion, session-184
SUB-SUBSYSTEMREQUIREMENTS-064	When a thruster firing is commanded, the catalyst bed heaters SHALL have maintained the catalyst bed temperature above 150 degrees Celsius for a minimum of 30 minutes prior to the first pulse to prevent cold-start degradation.	—	subsystem, propulsion, session-184
SUB-SUBSYSTEMREQUIREMENTS-065	The Propellant Management Assembly SHALL provide series-redundant propellant isolation via two normally-closed latch valves, each independently commandable, to prevent uncommanded propellant flow to the thrusters.	—	subsystem, propulsion, session-184
SUB-SUBSYSTEMREQUIREMENTS-066	The Thruster Assembly SHALL be qualified for a minimum of 50000 pulses and 10 hours cumulative firing time without degradation in thrust or specific impulse below specification limits.	—	subsystem, propulsion, session-184

Interface Requirements (IFC)

Ref	Requirement	V&V	Tags
IFC-INTERFACEDEFINITIONS-001	The interface between Telescope Assembly and Spectral Filter Mechanism SHALL present a collimated beam with wavefront error not exceeding lambda/10 RMS at 632.8 nm across the full field of view.	—	interface, optical-payload, session-177
IFC-INTERFACEDEFINITIONS-002	The interface between Focal Plane Array and Image Processing Unit SHALL use LVDS differential signalling at a data rate of 1.2 Gbps per channel with a maximum of 4 parallel channels.	—	interface, optical-payload, session-177
IFC-INTERFACEDEFINITIONS-003	The interface between Image Processing Unit and Onboard Data Handling Subsystem SHALL use SpaceWire links conforming to ECSS-E-ST-50-12C with a link rate of 200 Mbps per port.	—	interface, optical-payload, session-177
IFC-INTERFACEDEFINITIONS-004	The interface between Calibration System and Focal Plane Array SHALL provide a uniform illumination field with spatial non-uniformity of less than 1% across the detector active area during calibration sequences.	—	interface, optical-payload, session-177
IFC-INTERFACEDEFINITIONS-005	The interface between the Star Tracker Assembly and the AOCS Flight Software SHALL use a MIL-STD-1553B data bus transmitting attitude quaternion packets at 10 Hz with a maximum latency of 5 milliseconds from measurement to availability in the Kalman filter.	—	interface, aocs, session-178
IFC-INTERFACEDEFINITIONS-006	The interface between the AOCS Flight Software and the Reaction Wheel Assembly SHALL transmit torque command vectors at a minimum rate of 20 Hz via a CAN bus interface with command resolution of 0.001 Nm.	—	interface, aocs, session-178
IFC-INTERFACEDEFINITIONS-007	The interface between the GNSS Receiver and the Onboard Data Handling Subsystem SHALL provide a 1 PPS timing pulse with accuracy better than 100 ns and a serial NMEA/proprietary message stream at 1 Hz carrying position, velocity, and time data.	—	interface, aocs, session-178
IFC-INTERFACEDEFINITIONS-008	The interface between the Solar Array Assembly and the Power Control and Distribution Unit SHALL transfer unregulated DC power via redundant harness pairs rated for a maximum current of 40 A per string at voltages ranging from 60 V to 120 V across the full solar array operating temperature range of -150 C to +110 C.	—	interface, eps, session-179
IFC-INTERFACEDEFINITIONS-009	The interface between the Battery Assembly and the Power Control and Distribution Unit SHALL support bidirectional current flow of up to 30 A charge and 50 A discharge via a dedicated battery bus with continuous battery voltage, current, and temperature telemetry sampled at a minimum rate of 1 Hz.	—	interface, eps, session-179
IFC-INTERFACEDEFINITIONS-010	The interface between the Power Management Software and the Onboard Data Handling Subsystem SHALL use the spacecraft MIL-STD-1553B data bus for exchange of power telemetry packets, load shedding commands, and battery state of health parameters at a minimum update rate of 1 Hz.	—	interface, eps, session-179
IFC-INTERFACEDEFINITIONS-011	The interface between the Onboard Computer and the SpaceWire Network Router SHALL use ECSS-E-ST-50-12C SpaceWire protocol at 200 Mbps link rate with RMAP (Remote Memory Access Protocol) for register-level access to all connected nodes.	—	interface, obdh, session-180
IFC-INTERFACEDEFINITIONS-012	The interface between the SpaceWire Network Router and the Mass Memory Unit SHALL support sustained bulk data transfer at 800 Mbps using SpaceWire packet protocol with hardware flow control to prevent buffer overrun during continuous payload recording.	—	interface, obdh, session-180
IFC-INTERFACEDEFINITIONS-013	The interface between the SpaceWire Network Router and each Remote Terminal Unit SHALL operate at a minimum link rate of 10 Mbps using SpaceWire packet protocol, with each RTU addressable via a unique logical address for targeted housekeeping queries and command relay.	—	interface, obdh, session-180
IFC-INTERFACEDEFINITIONS-014	The interface between the S-Band Antenna Assembly and the S-Band Transponder SHALL use 50-ohm coaxial connections with SMA connectors, supporting the full 2025-2290 MHz band with insertion loss below 0.5 dB and VSWR below 1.3:1.	—	interface, ttc, session-182
IFC-INTERFACEDEFINITIONS-015	The interface between the S-Band Transponder and the Baseband Data Processor SHALL use a synchronous serial link at 1 Mbps for TM/TC baseband data exchange, with CCSDS transfer frame boundaries delimited by attached sync markers.	—	interface, ttc, session-182
IFC-INTERFACEDEFINITIONS-016	The interface between the Baseband Data Processor and the X-Band Transmitter SHALL use parallel LVDS signalling at a clock rate matching the selected symbol rate, with frame-aligned data valid assertion to synchronise modulator acquisition.	—	interface, ttc, session-182
IFC-INTERFACEDEFINITIONS-017	The interface between the X-Band Transmitter and the X-Band High-Gain Antenna SHALL use WR-112 rectangular waveguide with a maximum insertion loss of 0.3 dB across the 8.025-8.4 GHz band.	—	interface, ttc, session-182
IFC-INTERFACEDEFINITIONS-018	The interface between the Baseband Data Processor and the Mass Memory Unit SHALL use SpaceWire operating at 200 MHz link speed, supporting sustained payload data readout at 800 Mbps via dual redundant links with automatic failover.	—	interface, ttc, session-182
IFC-INTERFACEDEFINITIONS-019	The interface between the S-Band Transponder and the Onboard Computer SHALL use SpaceWire for telecommand packet delivery and telemetry source packet collection, with command latency from RF reception to OBDH delivery below 50 ms.	—	interface, ttc, session-182
IFC-INTERFACEDEFINITIONS-020	The interface between the Onboard Computer and the X-Band High-Gain Antenna gimbal controller SHALL use a CAN bus at 1 Mbps for antenna pointing commands, providing target azimuth and elevation angles at a minimum update rate of 10 Hz during ground station passes.	—	interface, ttc, session-182
IFC-INTERFACEDEFINITIONS-021	The interface between the Pulse-Tube Cryocooler and the Focal Plane Array SHALL use a flexible copper thermal strap with a conductance of at least 0.5 W/K to thermally isolate cryocooler vibrations from the detector mount.	—	interface, thermal-control, session-183
IFC-INTERFACEDEFINITIONS-022	The interface between the Thermostatically Controlled Heater System and the Power Control and Distribution Unit SHALL draw a maximum of 180 W from the unregulated 28 V bus for survival heaters during eclipse.	—	interface, thermal-control, session-183
IFC-INTERFACEDEFINITIONS-023	The interface between the Deployable Cryocooler Radiator and the Pulse-Tube Cryocooler SHALL use two flexible loop heat pipes with ammonia working fluid, each capable of transporting 50 W over a 1 m length.	—	interface, thermal-control, session-183
IFC-INTERFACEDEFINITIONS-024	The interface between the Heat Pipe Network and equipment mounting panels SHALL provide mechanical mounting provisions for at least 14 constant-conductance and 4 variable-conductance heat pipes with a total transport capacity exceeding 400 W.	—	interface, thermal-control, session-183
IFC-INTERFACEDEFINITIONS-025	The interface between the Launch Vehicle Adapter and the Primary Structure SHALL be a 1194 mm diameter bolt-circle with 24 M8 titanium bolts providing a minimum axial load capacity of 200 kN and a lateral load capacity of 75 kN.	—	interface, structure-mechanisms, session-184
IFC-INTERFACEDEFINITIONS-026	The interface between the Antenna Deployment Boom and the X-Band High-Gain Antenna gimbal SHALL provide a mechanical mounting flange with three-axis alignment accuracy of 0.1 degrees and a harness feedthrough for gimbal motor power and encoder signals.	—	interface, structure-mechanisms, session-184
IFC-INTERFACEDEFINITIONS-027	The interface between the Solar Array Deployment Mechanism and the Remote Terminal Unit SHALL provide discrete telemetry signals indicating hold-down status (stowed/released) and deployment latch engagement for each wing via redundant micro-switches.	—	interface, structure-mechanisms, session-184
IFC-INTERFACEDEFINITIONS-028	The interface between the Thruster Assembly and the AOCS Flight Software SHALL provide thruster on/off commanding via discrete pulse signals with a minimum pulse width of 20 ms and a maximum command latency of 10 ms from AOCS command issue to valve actuation.	—	interface, propulsion, session-184
IFC-INTERFACEDEFINITIONS-029	The interface between the Propellant Management Assembly and the Remote Terminal Unit SHALL provide analog telemetry of tank pressure (0-30 bar, 0.25% FSO accuracy), latch valve position status, and catalyst bed temperature for each thruster pair.	—	interface, propulsion, session-184
IFC-INTERFACEDEFINITIONS-030	The interface between the Propellant Tank and the Primary Structure SHALL provide a mounting bracket capable of supporting the filled tank mass of 53 kg under 8g axial and 3g lateral loads with thermal isolation standoffs limiting heat leak to less than 2 W.	—	interface, propulsion, session-184

Architecture Decisions (ARC)

Ref	Requirement	V&V	Tags
ARC-ARCHITECTUREDECISIONS-001	The Propulsion Subsystem SHALL use a monopropellant hydrazine blowdown system rather than bipropellant or electric propulsion. Rationale: hydrazine blowdown provides the simplest architecture for the 150 m/s delta-V budget, avoids oxidiser handling complexity, and offers heritage from ESA LEO missions. Electric propulsion was rejected due to the long thrust durations conflicting with imaging duty cycle requirements.	—	architecture, propulsion, session-186
ARC-ARCHITECTUREDECISIONS-002	The Onboard Data Handling Subsystem SHALL employ SpaceWire (ECSS-E-ST-50-12C) as the primary high-speed data bus interconnecting payload, mass memory, and platform subsystems. Rationale: SpaceWire provides deterministic latency, time-code distribution, and sufficient bandwidth (200 Mbps per link) for the 800 Mbps aggregate payload data rate when parallelised. MIL-STD-1553B was rejected for insufficient bandwidth; Ethernet-based alternatives lack sufficient spaceflight heritage for this mission class.	—	architecture, obdh, session-186
ARC-ARCHITECTUREDECISIONS-003	The Telemetry Tracking and Command Subsystem SHALL implement a dual-band architecture using S-band for telecommand uplink and housekeeping telemetry, and X-band for high-rate payload data downlink. Rationale: S-band provides reliable, near-omnidirectional coverage for commanding in all spacecraft attitudes including safe mode. X-band enables the 800 Mbps downlink rate required to empty the 2 Tbit mass memory within available ground station contact windows (typically 10 minutes per pass at 700 km altitude). A single-band approach cannot satisfy both the robust commanding and high-throughput data delivery requirements simultaneously.	—	architecture, ttc, session-186
ARC-ARCHITECTUREDECISIONS-004	The Thermal Control Subsystem SHALL use a pulse-tube cryocooler with a deployable radiator to cool the SWIR focal plane array, rather than passive radiative cooling or Stirling-cycle coolers. Rationale: the 150 K operating temperature required by the HgCdTe SWIR detector at 700 km sun-synchronous orbit cannot be reliably achieved by passive radiators alone due to Earth albedo and infrared flux. Pulse-tube cryocoolers provide vibration-free operation essential for sub-arcsecond pointing stability, unlike Stirling coolers which export significant mechanical disturbance at the drive frequency.	—	architecture, thermal-control, optical-payload, session-186
ARC-ARCHITECTUREDECISIONS-005	The Onboard Data Handling Subsystem SHALL employ a centralised radiation-hardened onboard computer with distributed remote terminal units for subsystem I/O acquisition, rather than a federated processing architecture. Rationale: a centralised OBC simplifies FDIR implementation and software qualification. Remote terminal units provide analogue and discrete I/O at each subsystem, reducing harness mass and EMI by keeping analogue signals short. The 400 MIPS processing budget is sufficient for all platform management tasks when payload image processing is handled by a dedicated IPU within the Optical Payload Subsystem.	—	architecture, obdh, session-186
ARC-ARCHITECTUREDECISIONS-006	The Attitude and Orbit Control Subsystem SHALL use a four-wheel reaction wheel assembly in a tetrahedral configuration with magnetorquer-based momentum unloading, rather than control moment gyroscopes or thruster-based attitude control. Rationale: reaction wheels provide the sub-arcsecond pointing stability and 0.001 deg/s stability required for 10 m GSD imaging. Magnetorquer unloading avoids propellant consumption for routine momentum management, preserving the hydrazine budget for orbit maintenance. CMGs were rejected as oversized for this satellite class. The tetrahedral four-wheel geometry provides single-fault tolerance against any one wheel failure.	—	architecture, aocs, session-186
ARC-ARCHITECTUREDECISIONS-007	The end-to-end data handling chain from payload to ground SHALL comply with CCSDS standards: CCSDS 122.0 for image compression, CCSDS AOS for transfer frame assembly, and CCSDS proximity-1 for command authentication. Rationale: CCSDS compliance ensures interoperability with existing ESA and partner ground station networks without bespoke protocol adapters. This architectural choice constrains the Baseband Data Processor and Image Processing Unit designs but eliminates ground segment development risk and enables multi-mission ground segment reuse.	—	architecture, obdh, ttc, session-186

Internal Diagrams

flowchart TB
  n0["component<br>AOCS Flight Software"]
  n1["component<br>Star Tracker Assembly"]
  n2["component<br>Reaction Wheel Assembly"]
  n3["component<br>Magnetorquer System"]
  n4["component<br>GNSS Receiver"]
  n5["component<br>AOCS Flight Software"]
  n6["component<br>Star Tracker Assembly"]
  n7["component<br>Reaction Wheel Assembly"]
  n8["component<br>Magnetorquer System"]
  n9["component<br>GNSS Receiver"]
  n6 -->|attitude quaternions| n5
  n9 -->|orbit PVT| n5
  n5 -->|torque commands| n7
  n5 -->|dipole commands| n8

Attitude and Orbit Control Subsystem — Internal

flowchart TB
  n0["system<br>Electrical Power Subsystem"]
  n1["component<br>Solar Array Assembly"]
  n2["component<br>Battery Assembly"]
  n3["component<br>Power Control and Distribution Unit"]
  n4["component<br>Solar Array Drive Mechanism"]
  n5["component<br>Power Management Software"]
  n1 -->|Unregulated DC power| n3
  n3 -->|Charge/discharge current| n2
  n5 -->|MPPT and load commands| n3
  n5 -->|Array pointing commands| n4
  n3 -->|28V regulated bus| n0

Electrical Power Subsystem — Internal

flowchart TB
  n0["component<br>Onboard Computer"]
  n1["component<br>Mass Memory Unit"]
  n2["component<br>SpaceWire Network Router"]
  n3["component<br>Remote Terminal Unit"]
  n4["component<br>OBDH Software"]
  n4 -->|Flight software executes on processor| n0
  n0 -->|Commands and telemetry packets| n2
  n2 -->|Payload and housekeeping data| n1
  n2 -->|Sensor queries and actuator commands| n3
  n0 -->|Store and retrieve commands| n1

Onboard Data Handling Subsystem — Internal

flowchart TB
  n0["component<br>S-Band Antenna Assembly"]
  n1["component<br>S-Band Transponder"]
  n2["component<br>Baseband Data Processor"]
  n3["component<br>X-Band Transmitter"]
  n4["component<br>X-Band High-Gain Antenna"]
  n5["external<br>Onboard Computer"]
  n6["external<br>Mass Memory Unit"]
  n7["actor<br>Ground Station"]
  n7 -->|S-band uplink/downlink RF| n0
  n0 -->|S-band RF via coax| n1
  n1 -->|TM/TC baseband data| n2
  n2 -->|CCSDS frames via LVDS| n3
  n3 -->|X-band RF via waveguide| n4
  n4 -->|X-band downlink RF| n7
  n6 -->|Payload data via SpaceWire| n2
  n5 -->|HK telemetry and scheduling| n2
  n1 -->|TC packets via SpaceWire| n5
  n5 -->|Antenna pointing commands| n4

Telemetry Tracking and Command Subsystem — Internal

flowchart TB
  n0["component<br>Pulse-Tube Cryocooler"]
  n1["component<br>MLI Blankets"]
  n2["component<br>OSR Radiator Panels"]
  n3["component<br>Heat Pipe Network"]
  n4["component<br>Heater System"]
  n5["component<br>Deployable Cryo Radiator"]
  n6["external<br>Focal Plane Array"]
  n7["external<br>PCDU"]
  n0 -->|150K cooling| n6
  n0 -->|waste heat| n5
  n3 -->|conducted heat| n2
  n4 -->|heater power| n7

Thermal Control Subsystem — Internal Block

Classified Entities

Entity	Hex Code	Description
Antenna Deployment Boom	CE951018	Deployable carbon-fibre boom for the X-band high-gain antenna on a LEO Earth observation satellite. Single-hinge deployment with torsion spring actuator and viscous fluid damper. Stowed length 0.4 m, deployed length 1.8 m. Carries the 0.6 m parabolic reflector and two-axis gimbal assembly (total tip mass 3.2 kg). Boom stiffness requirement: first deployed frequency >5 Hz to avoid coupling with AOCS control bandwidth. Hold-down via fusible wire with redundant heater elements. Includes thermal knife backup release. Must survive thermal cycling -120C to +100C without permanent deformation.
AOCS Flight Software	41F73B18	Attitude and orbit control flight software running on the AOCS processor within an Earth observation satellite. Implements attitude determination algorithms (extended Kalman filter fusing star tracker, gyro, and magnetometer data), guidance profiles for imaging/slew/safe modes, and closed-loop control laws commanding reaction wheels and magnetorquers. Manages mode transitions, fault detection, and orbit propagation.
Attitude and Orbit Control Subsystem	55F73218	Satellite pointing and orbit maintenance subsystem comprising reaction wheels, magnetorquers, star trackers, sun sensors, GPS receiver, and propulsion thrusters for precise attitude determination and control in low Earth orbit.
Baseband Data Processor	51F57008	Digital signal processing unit responsible for CCSDS transfer frame assembly, virtual channel multiplexing, forward error correction encoding (convolutional + Reed-Solomon for X-band, BCH for S-band), and modulation baseband signal generation. Receives payload data from the Mass Memory Unit via SpaceWire at rates up to 800 Mbps and housekeeping telemetry from the Onboard Computer. Implements AOS (Advanced Orbiting Systems) protocol stack per CCSDS 732.0. Manages downlink scheduling, prioritising real-time imaging data over stored data replay. FPGA-based architecture for deterministic throughput. Power consumption 25W.
Battery Assembly	D6D51218	Lithium-ion battery pack providing eclipse power for an earth observation satellite. Sized for 35-minute eclipse duration at full operational load. Includes cell modules, battery management electronics, thermal sensors, and heater circuits. Minimum 60000 charge-discharge cycles over 7-year mission life.
Calibration System	54F73218	Onboard radiometric and spectral calibration system for Earth observation satellite payload. Includes solar diffuser panel for absolute calibration and internal LED sources for relative calibration. Provides traceable reference measurements over 7-year mission lifetime.
Deployable Cryocooler Radiator	CE851018	Single-panel deployable radiator dedicated to pulse-tube cryocooler waste heat rejection. 0.8 m² white-painted aluminium honeycomb panel deployed by a hinge mechanism after launch. Rejection capacity: 100W at -40°C panel temperature, providing adequate margin for 75W EOL cryocooler waste heat. Connected to cryocooler compressor via two 1m flexible loop heat pipes with ammonia working fluid. Deployed panel has dedicated MLI backside insulation and sun shield.
Earth Observation Satellite	DDF75219	A low-Earth-orbit satellite system for remote sensing and Earth observation, carrying multispectral imaging payloads, with onboard data processing, attitude control, power generation, thermal management, and ground communication capabilities.
Electrical Power Subsystem	54D73218	Satellite power generation, storage, and distribution subsystem with deployable solar arrays, lithium-ion battery packs, power conditioning and distribution unit, and solar array drive mechanisms for continuous power supply through eclipse and sunlit orbital phases.
Focal Plane Array	D4D51018	Multispectral detector array for Earth observation satellite optical payload. Contains CCD and CMOS sensors covering visible, near-infrared, and shortwave infrared bands. Converts focused photons into electrical signals. Requires cryogenic cooling for SWIR channels.
Harness Assembly	C6851018	Complete satellite electrical harness for a 2000 kg LEO Earth observation satellite. Approximately 45 kg of cabling comprising power distribution harness (28V regulated bus, up to 80A peak), SpaceWire data harness (shielded twisted pair), pyrotechnic firing harness (isolated, shielded, with safe/arm switching), and RF coaxial cables (semi-rigid and flexible segments). Connectors are space-qualified micro-D and D-sub with EMI backshells. Harness routing designed for minimum EMI coupling between power, signal, and pyro circuits with >50mm separation and dedicated grounding scheme. All cables are radiation-resistant polyimide-insulated with ETFE jackets.
Heat Pipe Network	C6D51008	Aluminium-ammonia axially grooved heat pipe network distributing thermal loads between equipment mounting surfaces and radiator panels. 14 constant-conductance heat pipes (CCHP) embedded in honeycomb equipment panels, plus 4 variable-conductance heat pipes (VCHP) with nitrogen charge for temperature regulation during eclipse. Transport capacity: 20-40W per pipe over 0.5-1.2m lengths. Operating temperature range: -20°C to +50°C. Flight heritage: Sentinel-2 platform.
Image Processing Unit	50F53008	Onboard electronics unit for Earth observation satellite payload data processing. Performs radiometric correction, geometric correction, lossless and lossy compression, and formats data into CCSDS packets for downlink. Implements FPGA-based processing pipeline.
Launch Vehicle Adapter	CE851018	Clampband separation system interfacing the Earth observation satellite to the launch vehicle payload adapter. 1194 mm bolt-circle diameter (compatible with Vega-C Type-2 and Falcon 9 15-inch adapter). Marman clampband with pyrotechnic bolt cutters (redundant initiation) providing clean separation with tip-off rates below 1 deg/s. Separation springs sized for 0.5 m/s relative velocity. Includes separation confirmation micro-switches and umbilical disconnect for ground support equipment. Adapter cone is machined aluminium 7075-T6, mass 12 kg.
Magnetorquer System	54F53208	Three-axis magnetorquer system within an Earth observation satellite AOCS. Uses electromagnetic coils interacting with Earth's magnetic field to generate low-level control torques for momentum desaturation of reaction wheels. Also provides coarse attitude control during safe mode and initial acquisition. Driven by the AOCS flight software based on modelled geomagnetic field.
Mass Memory Unit	D4851018	Solid-state mass data recorder for an Earth observation satellite. Stores payload imagery, housekeeping telemetry, and science data between ground station passes. NAND flash based with EDAC protection, providing FIFO and random-access retrieval for downlink scheduling.
Multi-Layer Insulation Blankets	C6851018	Spacecraft thermal insulation consisting of 20-layer aluminised Kapton/Dacron MLI blankets covering all external surfaces except radiator panels and optical aperture. Effective emittance: 0.02. Total coverage area: ~12 m². Provides primary thermal isolation from solar flux and deep-space sink. Includes specialised cutouts and closeout details around harness penetrations, deployment mechanisms, and thruster nozzles. Grounding provisions for ESD protection.
Onboard Computer	D1F77018	Central flight processor for an Earth observation satellite data handling subsystem. Radiation-hardened processor executing flight software, managing command decoding, telemetry formatting, and autonomous operations scheduling. Interfaces with all subsystems via SpaceWire data bus.
Onboard Data Handling Software	41F77B18	Flight software for the onboard data handling subsystem of an Earth observation satellite. Manages command reception and validation, telemetry packet generation, autonomous fault detection isolation and recovery (FDIR), operations scheduling, and data routing between subsystems. Runs on the radiation-hardened onboard computer.
Onboard Data Handling Subsystem	51F77018	Satellite central computing and data management subsystem with radiation-hardened processors, solid-state mass memory for image storage, SpaceWire data bus, flight software, and command sequencing for autonomous operations and payload data compression.
Optical Payload Subsystem	D4C51018	Multispectral imaging instrument package for an Earth observation satellite, including visible, near-infrared, and shortwave infrared sensors, focal plane assemblies, and optical telescope assembly for capturing high-resolution Earth imagery.
Optical Solar Reflector Radiator Panels	C6851008	Four body-mounted OSR radiator panels providing passive heat rejection for spacecraft bus electronics. Total radiating area: 3.2 m² on anti-sun and Earth-facing panels. Silver-backed quartz OSR tiles with solar absorptance 0.08 BOL (0.15 EOL after 7 years UV degradation), IR emittance 0.80. Embedded aluminium-ammonia axially grooved heat pipes distribute thermal load across each panel. Rejects 600W orbit-average to deep space at panel temperature 0-20°C.
Power Control and Distribution Unit	D7F71008	Central power electronics unit for an earth observation satellite. Regulates unregulated solar array input to a 28V regulated bus. Performs maximum power point tracking, battery charge/discharge regulation, load switching via latching current limiters, and power telemetry acquisition. Radiation-hardened design for LEO environment.
Power Management Software	41B77B18	Flight software module executing on the onboard computer that controls the Electrical Power Subsystem of an earth observation satellite. Implements battery charge regulation algorithms, autonomous load shedding in under-power conditions, solar array MPPT control commands, fault detection and safe mode power configuration, and power telemetry formatting for downlink.
Primary Structure	CE851018	Honeycomb aluminium sandwich panel bus structure for a 2000 kg LEO Earth observation satellite. Box-frame configuration approximately 2.0m x 1.8m x 1.5m with 25mm thick Al-5056 honeycomb core panels bonded to 0.5mm CFRP facesheets on equipment-facing sides. Provides mounting surfaces for all subsystem electronics, payload, and propulsion equipment. Internal shear panels carry launch loads through to the launch vehicle adapter interface ring. Design loads: 8g axial, 3g lateral quasi-static with 1.25 safety factor. First natural frequency constraints: >25 Hz axial, >10 Hz lateral. Total structural mass budget 280 kg.
Propellant Management Assembly	D6851018	Fluidic control components for a monopropellant hydrazine propulsion system on a 2000 kg LEO satellite. Comprises two normally-closed latch valves (series-redundant for propellant isolation), a 10-micron sintered metal filter upstream of the thruster valves, fill-and-drain valve with capped service port, pressure transducer (0-30 bar, 0.25% FSO accuracy) on the pressurant side, and a bleed valve for ground servicing. All wetted materials compatible with hydrazine (Ti-6Al-4V, 304L stainless). Latch valves have independent heaters for thermal control during cold-case operations.
Propellant Tank	CE851018	Titanium diaphragm propellant tank for a monopropellant hydrazine system on a LEO Earth observation satellite. Ti-6Al-4V shell, 480mm diameter sphere, MEOP 24 bar with burst factor 2.0. Elastomeric diaphragm (AF-E-332) separating pressurant (GN2) from 45 kg hydrazine propellant. Expulsion efficiency >98%. Operating temperature range 10-50C maintained by spacecraft thermal control. Mass 8 kg empty. Tank mounted at spacecraft centre of mass on the Primary Structure central panel to minimise slosh-induced attitude disturbances.
Pulse-Tube Cryocooler	D6D51018	Thales LPT9310 single-stage pulse-tube cryocooler providing 1.5W cooling at 150K for HgCdTe SWIR focal plane detectors. Input power: 55W at beginning of life, 75W end of life. Compressor lifetime: >60,000 hours. Cold finger interfaces to detector cold plate via flexible copper thermal strap. Waste heat (55-75W) rejected via dedicated deployable radiator panel. Vibration export: <100mN at drive frequency, managed by active vibration cancellation using back-to-back compressor configuration.
Reaction Wheel Assembly	D6F51018	Four-wheel reaction wheel assembly in a redundant tetrahedral configuration within an Earth observation satellite AOCS. Provides fine three-axis torque for precision pointing and slew manoeuvres. Each wheel stores up to 50 Nms angular momentum, driven by brushless DC motors with closed-loop speed control. Used for agile imaging and maintaining pointing stability during payload operations.
Remote Terminal Unit	D5F57008	Distributed I/O module for an Earth observation satellite data handling subsystem. Acquires analog and digital sensor data from thermal, power, and structural health sensors, digitises readings, and transmits housekeeping telemetry packets to the onboard computer via SpaceWire. Also relays discrete commands to actuators and heaters.
S-Band Antenna Assembly	C6851018	Dual hemispherical S-band antenna assembly providing near-omnidirectional coverage for TM/TC links. Consists of two quadrifilar helix antennas mounted on opposing faces of the spacecraft bus, each with approximately 5 dBi peak gain and greater than -3 dBi over the hemisphere. Right-hand circular polarisation. Frequency range 2025-2290 MHz covering both uplink and downlink bands. Connected to the S-Band Transponder via coaxial cables and a hybrid coupler/diplexer network. Passive components with no active electronics — designed for 7+ year LEO lifetime without degradation.
S-Band Transponder	D4F57018	Full-duplex S-band transponder operating at 2025-2110 MHz uplink and 2200-2290 MHz downlink, providing CCSDS-compatible telecommand reception at up to 64 kbps and telemetry transmission at up to 512 kbps. Handles ranging tones for orbit determination. Implements coherent turnaround mode for two-way Doppler measurements. Interfaces with the Onboard Data Handling Subsystem via SpaceWire for TM/TC data exchange and with the RF distribution network via coaxial connectors. Power consumption approximately 15W in transmit mode. Radiation-hardened to 100 krad TID for 7-year LEO mission.
Solar Array Assembly	CEC51018	Deployable solar panel wings using GaAs triple-junction photovoltaic cells for an earth observation satellite. Generates 2500W+ end-of-life orbital average power in sun-synchronous LEO orbit. Includes substrate panels, cell strings, bypass diodes, and deployment hinges.
Solar Array Deployment Mechanism	CED51218	Hold-down and release mechanism for two deployable solar array wings on a LEO Earth observation satellite. Each wing is a 3-panel fold-out configuration using spring-driven hinges with viscous dampers for controlled deployment. Hold-down uses Frangibolt actuators (redundant pair per wing) rated at 22 kN. Deployment sequence initiated by OBDH command, completes within 120 seconds. Must survive launch vibration without premature release (positive margin against 3-sigma random vibe). Includes deployment status micro-switches for telemetry confirmation. Total mechanism mass per wing: 4.5 kg.
Solar Array Drive Mechanism	DFF71208	Single-axis rotary drive mechanism that orients solar array wings toward the Sun to maximise power generation on an earth observation satellite. Includes stepper motor, harmonic drive gearbox, slip ring assembly for power transfer, and angular position encoder. Must operate continuously for 7+ years in LEO thermal cycling environment.
SpaceWire Network Router	D4E57018	High-speed data network router for an Earth observation satellite onboard data handling subsystem. Provides deterministic packet routing between onboard computer, payload instruments, attitude sensors, and telemetry subsystem via SpaceWire protocol. Supports time-code distribution for synchronisation.
Spectral Filter Mechanism	D6951008	Electromechanical filter selection mechanism for Earth observation satellite multispectral imager. Contains dichroic beam splitters and bandpass filters to separate incoming radiation into discrete spectral channels. Stepper motor driven filter wheel with position encoding.
Star Tracker Assembly	D5F73218	Optical star tracker sensor unit within an Earth observation satellite AOCS. Uses CCD or CMOS detector to image star fields, performs on-chip star pattern matching against an onboard catalogue, and outputs inertial attitude quaternions at 10 Hz with arcsecond-class accuracy. Operates autonomously after initial acquisition.
Structure and Mechanisms Subsystem	CE851018	Primary load-bearing and deployment subsystem for a 2000 kg-class LEO Earth observation satellite in sun-synchronous orbit at 700 km altitude. Comprises the central bus structure (honeycomb aluminium panels forming a box configuration), solar array deployment and hold-down mechanisms, X-band high-gain antenna deployment boom and gimbal mount, launch vehicle adapter ring (1194 mm diameter compatible with Vega-C or Falcon 9 ESPA), and the satellite separation system. Must withstand quasi-static launch loads of 8g axial and 3g lateral, acoustic environment up to 140 dB OASPL, and provide mounting interfaces for all subsystem equipment with thermal and electrical isolation where required. First natural frequency above 25 Hz axial and 10 Hz lateral.
Telemetry Tracking and Command Subsystem	54E57018	Satellite communication subsystem for ground segment link, including S-band transponders for telemetry and command, X-band transmitter for high-rate payload data downlink, antennas, and RF switching network.
Telescope Assembly	CE851018	Primary and secondary mirror assembly with focal optics for an Earth observation satellite. Collects and focuses incoming electromagnetic radiation onto the focal plane detector array. Korsch three-mirror anastigmat design providing wide field of view with minimal optical aberration.
Thermal Control Subsystem	51D71208	Satellite thermal management subsystem using passive techniques (multilayer insulation, radiators, heat pipes, thermal coatings) and active heaters to maintain all components within operational temperature limits through extreme orbital thermal cycling between sunlit and eclipse phases.
Thermostatically Controlled Heater System	55F73008	Distributed heater system with 48 Kapton-film heaters (2W to 20W each) and 96 platinum resistance temperature sensors (Pt1000). Managed by a dedicated heater controller unit with 32 thermostat channels providing proportional-integral control. Total survival heater power budget: 180W during eclipse. Protects batteries (>0°C), propellant lines (>5°C), and star trackers (>-30°C) during eclipse and safe-mode. Autonomous operation independent of OBC for survival heaters.
Thruster Assembly	D6C53018	Four 1N monopropellant hydrazine thrusters (Shell 401 catalyst) for orbit maintenance and deorbit of a 2000 kg LEO Earth observation satellite. Two redundant pairs mounted on the anti-velocity face, canted 10 degrees off-axis to balance thrust torques. Each thruster provides 0.9-1.1N thrust over the blowdown pressure range (22-5.5 bar), specific impulse 220s nominal. Minimum impulse bit 50 mN-s for fine orbit adjust. Catalyst bed heaters (redundant, 10W each) maintain bed temperature above 150C. Valve response time <20ms. Qualified for 50000 pulses and 10 hours cumulative firing.
X-Band High-Gain Antenna	DEC51018	Steerable X-band parabolic reflector antenna with 0.5m aperture diameter providing 34 dBi gain at 8.2 GHz for high-rate payload data downlink. Two-axis gimbal mechanism allows ±90 degree steering to track ground stations during LEO passes. Pointing accuracy better than 0.3 degrees. Right-hand circular polarisation with axial ratio below 1.5 dB. Connected to the X-Band Transmitter via waveguide. Gimbal motors and angle encoders interface with the Onboard Data Handling Subsystem for antenna pointing control during ground station passes. Mass approximately 8 kg including gimbal.
X-Band Transmitter	D4E55018	High-power X-band transmitter operating at 8.025-8.4 GHz for payload data downlink. Supports selectable data rates from 150 Mbps to 800 Mbps using OQPSK modulation with CCSDS convolutional and Reed-Solomon coding. Output RF power of 8W at antenna port. Accepts formatted CCSDS transfer frames from the Baseband Data Processor via parallel LVDS interface. Includes solid-state power amplifier (SSPA), upconverter, and local oscillator. Power consumption 45W nominal. Designed for LEO Earth observation missions with typical ground station pass durations of 8-12 minutes.

Decomposition Relationships

Part-Of

Component	Belongs To
Optical Payload Subsystem	Earth Observation Satellite
Attitude and Orbit Control Subsystem	Earth Observation Satellite
Electrical Power Subsystem	Earth Observation Satellite
Telemetry Tracking and Command Subsystem	Earth Observation Satellite
Onboard Data Handling Subsystem	Earth Observation Satellite
Thermal Control Subsystem	Earth Observation Satellite
Telescope Assembly	Optical Payload Subsystem
Focal Plane Array	Optical Payload Subsystem
Image Processing Unit	Optical Payload Subsystem
Spectral Filter Mechanism	Optical Payload Subsystem
Calibration System	Optical Payload Subsystem
Star Tracker Assembly	Attitude and Orbit Control Subsystem
Reaction Wheel Assembly	Attitude and Orbit Control Subsystem
Magnetorquer System	Attitude and Orbit Control Subsystem
GNSS Receiver	Attitude and Orbit Control Subsystem
AOCS Flight Software	Attitude and Orbit Control Subsystem
Solar Array Assembly	Electrical Power Subsystem
Battery Assembly	Electrical Power Subsystem
Power Control and Distribution Unit	Electrical Power Subsystem
Solar Array Drive Mechanism	Electrical Power Subsystem
Power Management Software	Electrical Power Subsystem
Onboard Computer	Onboard Data Handling Subsystem
Mass Memory Unit	Onboard Data Handling Subsystem
SpaceWire Network Router	Onboard Data Handling Subsystem
Remote Terminal Unit	Onboard Data Handling Subsystem
Onboard Data Handling Software	Onboard Data Handling Subsystem
S-Band Transponder	Telemetry Tracking and Command Subsystem
X-Band Transmitter	Telemetry Tracking and Command Subsystem
S-Band Antenna Assembly	Telemetry Tracking and Command Subsystem
X-Band High-Gain Antenna	Telemetry Tracking and Command Subsystem
Baseband Data Processor	Telemetry Tracking and Command Subsystem
Pulse-Tube Cryocooler	Thermal Control Subsystem
Multi-Layer Insulation Blankets	Thermal Control Subsystem
Optical Solar Reflector Radiator Panels	Thermal Control Subsystem
Heat Pipe Network	Thermal Control Subsystem
Thermostatically Controlled Heater System	Thermal Control Subsystem
Deployable Cryocooler Radiator	Thermal Control Subsystem
Structure and Mechanisms Subsystem	Earth Observation Satellite
Primary Structure	Structure and Mechanisms Subsystem
Solar Array Deployment Mechanism	Structure and Mechanisms Subsystem
Launch Vehicle Adapter	Structure and Mechanisms Subsystem
Antenna Deployment Boom	Structure and Mechanisms Subsystem
Harness Assembly	Structure and Mechanisms Subsystem
Thruster Assembly	Propulsion Subsystem
Propellant Tank	Propulsion Subsystem
Propellant Management Assembly	Propulsion Subsystem

Connections

From	To
Telescope Assembly	Spectral Filter Mechanism
Spectral Filter Mechanism	Focal Plane Array
Focal Plane Array	Image Processing Unit
Calibration System	Focal Plane Array
Star Tracker Assembly	AOCS Flight Software
GNSS Receiver	AOCS Flight Software
AOCS Flight Software	Reaction Wheel Assembly
AOCS Flight Software	Magnetorquer System
Solar Array Assembly	Power Control and Distribution Unit
Battery Assembly	Power Control and Distribution Unit
Power Management Software	Power Control and Distribution Unit
Power Management Software	Solar Array Drive Mechanism
Onboard Computer	SpaceWire Network Router
Onboard Computer	Mass Memory Unit
SpaceWire Network Router	Remote Terminal Unit
Onboard Data Handling Software	Onboard Computer
SpaceWire Network Router	Mass Memory Unit
S-Band Antenna Assembly	S-Band Transponder
S-Band Transponder	Baseband Data Processor
Baseband Data Processor	X-Band Transmitter
X-Band Transmitter	X-Band High-Gain Antenna
Baseband Data Processor	Mass Memory Unit
Baseband Data Processor	Onboard Computer
S-Band Transponder	Onboard Computer
X-Band High-Gain Antenna	Onboard Computer
Image Processing Unit	Mass Memory Unit
Pulse-Tube Cryocooler	Focal Plane Array
Pulse-Tube Cryocooler	Deployable Cryocooler Radiator
Heat Pipe Network	Optical Solar Reflector Radiator Panels
Thermostatically Controlled Heater System	Power Control and Distribution Unit
Launch Vehicle Adapter	Primary Structure
Solar Array Deployment Mechanism	Primary Structure
Antenna Deployment Boom	X-Band High-Gain Antenna
Primary Structure	Optical Payload Subsystem
Harness Assembly	Power Control and Distribution Unit
Harness Assembly	SpaceWire Network Router
Propellant Tank	Propellant Management Assembly
Propellant Management Assembly	Thruster Assembly
Thruster Assembly	Primary Structure
Propellant Tank	Primary Structure
Thruster Assembly	AOCS Flight Software
Propellant Management Assembly	Remote Terminal Unit
Solar Array Assembly	Solar Array Deployment Mechanism
Solar Array Assembly	Solar Array Drive Mechanism

Produces

Component	Output
Telescope Assembly	focused optical beam
Focal Plane Array	raw digital image data
Image Processing Unit	compressed CCSDS image packets
Spectral Filter Mechanism	spectrally filtered radiation
Calibration System	radiometric calibration reference
Star Tracker Assembly	inertial attitude quaternions
Reaction Wheel Assembly	three-axis control torque
Magnetorquer System	magnetic desaturation torque
GNSS Receiver	orbit position-velocity-time solution
AOCS Flight Software	actuator command profiles
Solar Array Assembly	unregulated DC power from photovoltaic conversion
Battery Assembly	regulated eclipse power and transient load buffering
Power Control and Distribution Unit	28V regulated bus power and load switching commands
Solar Array Drive Mechanism	optimal solar array sun-pointing orientation
Power Management Software	charge control commands and load shedding decisions
Onboard Computer	command execution and telemetry formatting
Mass Memory Unit	stored payload and housekeeping data for downlink
SpaceWire Network Router	deterministic packet routing and time-code distribution
Remote Terminal Unit	digitised sensor telemetry and discrete actuator commands
Onboard Data Handling Software	FDIR decisions, operations schedules, and command validation
S-Band Transponder	CCSDS telecommand packets and ranging turnaround signals
X-Band Transmitter	modulated X-band RF carrier for payload data downlink
S-Band Antenna Assembly	omnidirectional S-band RF coverage for TM/TC
X-Band High-Gain Antenna	high-gain directed X-band beam for payload downlink
Baseband Data Processor	FEC-encoded CCSDS transfer frames and modulation baseband
Telescope Assembly	diffraction-limited image at focal plane
Focal Plane Array	raw multi-spectral pushbroom pixel data at 2.4 Gbps
Spectral Filter Mechanism	spectrally separated VNIR and SWIR optical channels
Image Processing Unit	radiometrically corrected compressed image segments
Calibration System	calibration reference illumination and dark frames
Pulse-Tube Cryocooler	150K cold-tip temperature for SWIR detector cooling
Multi-Layer Insulation Blankets	passive thermal isolation from external environment
Optical Solar Reflector Radiator Panels	passive heat rejection of 600W bus electronics dissipation
Heat Pipe Network	isothermal equipment panel temperature distribution
Thermostatically Controlled Heater System	survival and operational temperature maintenance during eclipse
Deployable Cryocooler Radiator	100W cryocooler waste heat rejection at -40C
Primary Structure	structural load path from all equipment to launch vehicle interface
Solar Array Deployment Mechanism	deployed solar array wings in operational configuration
Launch Vehicle Adapter	mechanical and electrical interface to launch vehicle
Antenna Deployment Boom	deployed antenna platform with required stiffness and alignment
Harness Assembly	electrical power and signal interconnection between all subsystems
Thruster Assembly	thrust for orbit maintenance and deorbit manoeuvres
Propellant Tank	pressurised hydrazine propellant supply to thrusters
Propellant Management Assembly	regulated propellant flow and system health telemetry
Solar Array Assembly	raw photovoltaic electrical power at 60-100V DC