Satellite System (SAT)
JPL leads the development of the Satellite System (SAT) in partnership (contract) with Astrium GmbH. Astrium provides major elements of two flight satellites based on an existing small satellite design for the CHAMP, GRACE and SWARM missions. The GRACE-FO satellites are shown in the following figure.
The Satellite System (SAT) consists of the following sub-systems where most are available with main and redundant units.
Telemetry, Tracking & Control (TT&C)
The Telemetry, Tracking & Control (TT&C) activities are carried out using a pyro-deployed S-Band receive and transmit antenna, mounted on a nadir-facing deployable boom. Two back-up zenith antennae, one each for transmitting and receiving, along with the appropriate RF electronics assembly, complete the telemetry and telecommand sub-system.
The telecommand function of the satellite is designed according to the ESA CCSDS (Consultative Committee for Space Data Systems) Packet Telecommand Standard tailored for GRACE-FO Packet Utilization Standard (G-PUS) with adaptations mutually agreed with the German Space Operation Center (GSOC). The satellites support the command and control capabilities of the MOS by means of:
- high priority commands of priority 1 (HPC1), which are directly handled by the telecommand decoder and by-passes all OBC (On-board Computer) software
- normal telecommands, which will be processed by the OBC on-board software.
After reception of the uplinked command stream via the S-band antenna and the receiver of the RF Electronic Assembly, the telecommands are decoded in above two command categories. The High Priority Commands of level 1 (HPC 1) are directly executed within the telecommand module of the OBC; i.e. corresponding bi-stable relays are set. The normal telecommands are read from the telecommand handler of the on-board flight S/W via system calls. The telecommand handler further validates and converts the telecommand packets into on-board command packets (OCP's). The on-board command packets are further distributed according to their indicated functionality.
The Power System is responsible for generation, storage, conditioning and distribution of electrical power in accordance with instrument and satellite bus users needs. Electrical energy is generated using solar arrays of triple junction Gallium Arsenid (GaAs) cells, placed on the top and side exterior surface of the satellites. Excess energy is stored in a battery of Li-Ion cells with a capacity of 66 Ah at mission start. The power bus delivers unregulated power to all users at the respective user interface.
The Thermal Control System consists of 96 independent heater circuits, 128 YSI-type thermistors and 36 PT-type thermistors for in-flight temperature housekeeping, monitoring and heater control, as well as for on-ground verification testing.
On-Board Computer (OBC) System
The On-Board Computer (OBC) System provides processor and software resources, as well as necessary I/O capabilities for AOCS (Attitude and Orbit Control System), Power and Thermal Systems operations, including necessary fault detection, isolation and recovery operations.
Attitude and Orbit Control System (AOCS) The Attitude and Orbit Control System (AOCS) consists of sensors, actuators and software to:
- Provide adequate knowledge of satellite attitude during all phases of the mission,
- Generate on-board error signals to accurately maintain satellite attitude,
- Provide necessary orbital control to satisfy the GRACE-FO mission requirements.
The sensors include a Coarse Earth Sun Sensor (CESS), an Inertial Measurement Unit (IMU) and a fluxgate magnetometer, as well as the Star Tracker Assembly (STR) and GNSS receiver described in the Satellite Instrument System (SIS) page.
The CESS provides for omni-directional, coarse attitude measurement in the initial acquisition, survival and stand-by modes of the satellite. It comprises of six thermistors orthogonally mounted on the satellite. By assuming that the Sun is the hottest object in the field of view and the earth is the second hottest object in the field of view, the CESS provides the Sun and Earth vectors relative to the body frame at a rate of 1 Hz and accuracy of 5-degrees [TBC].
The IMU is used in survival modes and provides 4-axis rate information. The unit comprises of three solid-state fiber optic gyros, and three solid-state silicon accelerometers that measures velocity and angle changes in a coordinate system fixed relative to its case.
A fluxgate magnetometer provides additional rate information.
The actuators for the AOCS include a Cold Gas Assembly and a Magnetorquer System. The GN2 reaction control system includes two pressure vessels, valves, regulators and filters, along with 12 attitude control thrusters and two orbit control thrusters. Three magnetorquers with linear dipole moments of 27.5 Am2 complete the set of AOCS actuators.
Star Tracker Assembly (STR)
The µASC Star Tracker Assembly (STR) determines the orientation of the satellite by tracking it relative to the position of the stars. These measurements are used for fine-pointing and on-board control of the satellite. Additionally they are required for the interpretation of measurements made in the satellite reference frame, such as those from the SuperSTAR accelerometer (see in the Satellite Instrument System (SIS) page).
The SCA consists of three temperature controlled CCD star cameras mounted to the accelerometer, along with the respective baffle assemblies. The STR delivers its video frames to the On-Board Computer (OBC), which then computes the attitude quaternions. The OBC also acts as the power and command/control interface to the STR. Once switched on and initialized, the STR proceeds with automatic coarse attitude acquisition and then on to fine attitude derivation.
Centre of Mass Trim (CMT) Assembly
The Center of Mass Trim (CMT) Assembly consists of six (two per axis) Mass Trim Mechanisms (MTM), associated electronics, and the power and signal harness. Each MTM consists of a trim mass driven on a nut rotor with a stepper motor. The CMT Assembly is used to center the center of gravity (CG) of the satellite at the center of the proof-mass of the accelerometer after CG calibration maneuvers.
Die Entwicklung der Satelliten und seiner Komponenten (SAT) wird von JPL geleitet und in enger Partnerschaft auf Vertragsbasis durch Astrium GmbH realisiert. Die wesentlichen Komponenten der beiden Satelliten basieren dabei auf Erfahrungen die beim Bau der Kleinsatelliten wie CHAMP, GRACE oder SWARM bei Astrium gemacht wurden. Die GRACE-FO Satelliten sind im nachfolgenden Bild dargestellt. Details zur Datenübertragung, Lageregelung, Temperaturkontrolle oder Stromversorgung sind in den entsprechenden englischen Seiten zu finden.