Fast Auroral SnapshoT Explorer
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The FAST Spacecraft


For a nice diagram of the spacecraft and descriptions of its instruments, please see the Education Page.


The following is quoted from Pfaff et. al., An Overview of the Fast Auroral SnapshoT (FAST) Mission.

Structure

The FAST primary structure is constructed of aluminum and includes a single deck on which the instruments and electronic boxes are mounted as shown in the exploded drawing in Figure 5. There are two magnetometer booms located 180 degrees relative to each other in the spin plane, which are stowed along the spin axis for launch. In addition, there are two axial stacer electric field booms and four radial, equally spaced, wire electric field booms that are deployed after launch.

The instrument deck layout is shown Figure 6. Placing all of the instruments and main spacecraft components on a single deck whose plane includes the spin plane electric field wire booms and deployed magnetometer booms is an important design feature of the FAST spacecraft. Such a single plane design, together with the placement of the instruments at the furthest edges of this plane, optimizes the spacecraft moment of inertia such that the spacecraft is thus able to support the longest possible spin axis electric field booms [see Pankow et al., Deployment Mechanisms on the FAST Satellite: Radial Wires, Stiff Axials, and Magnetometer Booms]. Notice also in Figure 6 that the design facilitates the fact that all connectors are exposed and accessible for testing purposes when all of the instruments and the MUE are installed on the single deck.

Above and below the instrument deck, the solar array attaches to an essentially hollow "shell" whose main purpose is to maximize the solar array area within the limits of the Pegasus shroud. The relatively low mass of the solar array shell has minimum impact to the satellite moment of inertia. . . .The result is a satellite whose shape is optimized for power (e.g., solar cell area) and whose moment of inertia is optimized to support the longest possible spin axis electric field booms.


Power

The FAST power system is a direct energy transfer system. Excess charge is dissipated in shunts to prevent battery overcharge. Shunt driver boxes external to the Mission Unique Electronics (MUE) are provided to switch the excess solar array current to the shunts. The battery on the FAST spacecraft is a 9 Ampere-hour super NiCad battery. The FAST power system is described in detail by Schnurr et al. [1995].

The MUE power circuitry provides voltage/temperature (V/T) control, current control, over-voltage control and a precision current monitor that is used as the sensor for the software Amp-Hour integrator. The battery charge control circuitry holds eight relays which are used to enable and configure the state of the controller. All relays are magnetic latching to hold states in between modes and commanded configuration changes. If the voltage is out of tolerance for three cycles, the MUE software will shut down non-essential loads. The charge control card drives the shunt regulator. The voltage and temperature controls are monitored and controlled according to pre-determined voltage-temperature (V/T) levels. In V/T control mode, when the battery reaches the selected voltage limit, the V/T controller will activate the shunt regulator. The battery temperature is read from platinum wire temperature sensors located in the battery. Sixteen different V/T levels are available and commandable from the ground. The overvoltage controller activates the shunt regulators when the battery approaches 34.5 +/- 0.5 VDC.

The power distribution function provides 11 relays to apply unregulated +31-32 Volt power to the instruments, deployment mechanisms, heaters, and ACS sensors. Bus power to all devices except the transponder is switched by the MUE. The transponder cycles power to the transmitter section only when the transmitter is in use.


Solar Array

A body-mounted solar array was required by the FAST science instruments in order to minimize plasma disturbances (both for field and particle measurements) around the spacecraft. Furthermore, to minimize electrostatic charging and stray magnetic fields, the mission required a conducting solar array (i.e., sealed inside a Faraday box) with circuitry to cancel magnetic fields generated by the solar array strings and wiring.

The FAST solar array uses 8,225 solar cells that comprise a total area of 2.58 m2. The cells are 0.14 mm thick, 18.5% efficient single junction GaAs/Ge. A 60-mil thick coverglass was included to protect the cells from the harsh radiation environment. This coverglass is composed of fused silica with the standard space qualified anti-reflective front surface and ultraviolet reflective rear surface coatings and an additional outer surface transparent conductive coating of indium tin oxide (ITO) to act as part of the Faraday box. Details of the FAST solar array including a discussion of the solar cells with connection "V-clips," boom shadowing, and other features of the array design are discussed in Kruer and Lyons [1994].

During its first 3.5 years in space, the FAST solar array output has varied from 60-130 W depending on beta angle (the angle between the orbital plane and the sun-earth direction), as shown in Figure 8(a) [Lyons, personal communication, 2000]. Figure 8(b) shows the projected area of the solar array averaged over a spin period for various beta angles. Notice that the measured solar array power closely tracks this projected area. The available solar array power degrades with time, at a rate of about 10% over 3 years, as shown in Figure 8(a). Throughout this period, the average spacecraft bus voltage has been maintained near +31.4 V. The excellent solar array performance is somewhat better than predicted as the FAST orbit exposes the satellite to an intense radiation environment combined with high operating temperatures and significant atomic oxygen flux at perigee.


Thermal

Temperatures are controlled primarily by passive thermal control elements. As described in Parrish [1999], only the battery has a dedicated radiator and thermostatically controlled heaters. (The transponder uses heaters during safehold operations). All of the main electronics are heat sunk to the main deck. The body mounted solar array is conductively and radiatively coupled to the equipment deck and equipment cavity. The equipment power is transferred to the body array where it is then radiated to space. When the sun vector is nearly perpendicular to the spacecraft's spin axis, the body array provides a nearly room temperature environment for the electronic equipment.

Multi-layer insulation is positioned on the interior of select areas of the array to trim and adjust the equipment deck temperatures. Heat is also transferred to and from the thrust cone which supports the equipment deck. The thrust cone attaches to the launch vehicle irridited aluminum interface ring which protrudes from the bottom of the spacecraft.

During its first three years, the FAST internal instrument disk temperatures varied between -10 and 35 degrees C. The spacecraft has, in general, been somewhat warmer than predicted, but is still within temperature limits, as discussed by Parrish [1999]. In fact, the warmer spacecraft allowed data to be gathered during times when it was originally believed that sun-avoidance attitude maneuvers would be needed to avoid the cold temperatures that were predicted prior to launch.


Mission Unique Electronics

The spacecraft command and data handling (C&DH) system is embedded within the spacecraft electronics module. This module is known as the Mission Unique Electronics (MUE), and uses a pair of 2 MHz 80C85 8-bit microprocessors, with 72 kbyte ROM and 320 kbyte RAM. A block diagram of the MUE is shown in Figure 9. The MUE performs telecommand reception, stored command processing, telemetry data collection and generation, data encoding and decoding attitude control, power management and battery charge control, launch vehicle interface information and basic spacecraft health and safety functions. Also within the MUE is an 800 kbyte RAM recorder for capturing spacecraft safing events.


Spacecraft (MUE) Software

The FAST spacecraft (MUE) software is partitioned into three main subsystems: C&DH, ACS, and Power Management. A custom Operating System (OS) provides software task control and basic Input/Output (I/O) functions to the subsystems. The software processing is divided between the two 8085 microprocessors, with most C&DH functions in one, and the Power and ACS subsystems in the other. All of the software subsystems operate under three modes: Launch, Initial Acquisition, and Normal modes. The spacecraft was switched to normal mode after initial acquisition, where it will remain for the duration of the mission.

The C&DH command uplink system processes CCSDS formatted ground or stored commands and distributes them to other subsystems including the IDPU. The command system can process two types of stored commands: Absolute Time Commands (ATC), and Relative Time Commands (RTC). ATCs are stored in either of two buffers holding up to 512 commands each. These buffers hold commands to control spacecraft and some instrument operations that must be performed on a fixed timeline. RTCs can be stored in any of 64 sequence buffers of 13 commands each. Once commanded to start, these commands execute at times relative to each other. Up to 16 of these RTC buffers may be active at any time.

The power management subsystem software consists of two sections: the command processor and the Amp-Hour Integrator, both executing at 1 Hz. The Amp-Hour Integrator computes the battery state of charge based on the battery current and temperature. When the charge reaches 100%, the software issues a trickle charge command. The state of charge variable and several other variables computed by the power subsystem are also used by ground operators and safing checks for spacecraft power management. The command processing section interprets and executes ground commands to control spacecraft power distribution.


Instrument Data Processing Unit

The instrument data processing unit (IDPU), described [in Harvey et al., The FAST Spacecraft Instrument Data Processing Unit], manages and controls the multiple functions of the fields and particles instruments, including boom deployments. Within the IDPU is a high density, 1 Gbit solid state recorder, which includes a 1-2 Mbyte partition for spacecraft health and safety data. The MUE and IDPU communicate via a simple serial interface at 9600 baud, with a high rate interface for the IDPU to directly access the transponder for downlink of science data. The IDPU contains a 32-bit microprocessor.


Attitude Control System

The primary objectives of the Attitude Control System (ACS) are to provide autonomous spin and precession control following separation from the launch vehicle, spacecraft spin and precession control during the normal mode operations to meet the science imposed attitude requirements, and maintain a spin rate and spin axis attitude consistent with the power and thermal requirements. The ACS is also responsible for processing ACS related ground commands and providing telemetry associated with the ACS.

The ACS for FAST is designed to maintain the spacecraft attitude as a simple spinner with a rotation rate of approximately 12 rpm. Pointing requirements are met by utilizing a complement of sensors, torquers and standard electronics in the MUE. Spin rate and spin axis orientation are maintained using two magnetic torquer coils. Either one spinning sun sensor, one horizon crossing indicator, or the spacecraft magnetometer can be used to measure the actual spin period. In addition to the active control elements, a fluid ring damper provides passive nutation control. The ACS provides closed-loop spin-rate control. Spin axis precession is performed open loop and is closed via ground commands. The near orbit-normal spinner uses electromagnets to keep up with the daily orbit precession and perform sun angle avoidance maneuvers to maintain the sun angle to less than 60 degrees, if necessary.


Attitude Knowledge

FAST attitude knowledge consists of two elements: spin axis pointing and spin phase. Spin-axis pointing is verified by orbit detrending between the model magnetic field and the observed magnetic field. Nominal accuracy using actual FAST data is typically 0.1 degrees, although this is a factor of two worse on orbits around torquer coil operations.

The spin-phase error is much less than 0.1 degrees in sunlight (sun-sensor determined), but jitter of this order is present when FAST is in eclipse (where the equivalent sun-phase is determined from using nadir-phase information). There are also frequent phase skips at eclipse entry and exit. In other words, although the spin period may be reasonably accurate in eclipse, the absolute phase can have problems.

An additional source of attitude error derives from spacecraft nutation. This error can be quite large, ~0.2 degrees with a 30 second period. Such effects are largest immediately after torquer operations. After computation on the ground, the FAST attitude knowledge is accurate to within approximately 0.1 degree.


Telemetry Rates and Transmitter

FAST uses a standard 5-Watt NASA transponder that receives commands at a data rate of 2 kbps that are transmitted as non-return to zero (NRZ) bi-phase modulation on a 16 kHz sub-carrier at 2.03964 GHz. Telemetry data is transmitted at 4 digital data rates (4 kbps, 900 kbps, 1.5 Mbps, and 2.25 Mbps) using NRZ phase modulation directly on the carrier. A multi-element micropatch antenna mounted on a boom above the spacecraft is used to support ground communications.


Radiation

The high radiation environment of FAST demanded careful part selection and careful attention to the structure in the design phase. All electronic components are supported by a machined aluminum equipment deck with an aluminum honeycomb radiation shield cover above. The solar array substrate is also honeycomb and provides radiation shielding on the sides. The solar cells are covered with extra-thick coverglass to help reduce the total-dose of radiation seen by the solar cells. Samples from each lot of parts were tested for single event effects and total dose effects. One reason why 8085 processors were selected for the MUE and the 32C016 processor was selected for the IDPU is that these were available as radiation-hard parts.


Electric and Magnetic Shielding

In order to minimize plasma disturbance around the spacecraft, the instruments require a solar array with magnetic and electrostatic cleanliness values of 7.85 gammas and 0.1V maximum voltage differential across the spacecraft outer surface, both an order of magnitude less than required for any array previously built by NASA. Throughout the design and fabrication phase, care was taken to reduce the electric and magnetic fields produced by the spacecraft so that the scientific measurements of these geophysical quantities would not be corrupted. To this end, the FAST solar array incorporates an integral Faraday cage. Each coverglass is coated with conductive (and transparent) indium tin oxide (ITO). The gaps between coverglass are filled V-shaped pieces of metal which are attached to the coverglass with palladium-filled epoxy. The edges of the solar arrays are covered with conductive foil. All external conductive surfaces are electrically connected together and connected to the aluminum structure, presenting an equipotential surface to the space plasma environment. The cells and backwiring were laid out to minimize magnetic fields. All harnessing which carries appreciable current is twisted to its return. Where redundant power lines were required, quad- and hex-twisted shielded lines were used so that all current would be close to its return. Power distribution relays were oriented to cancel internal magnetic fields. A loop was added to the battery harness to cancel magnetic fields produced within the battery. The spacecraft produces no more than 1 nT of DC magnetic fluctuations at the fluxgate magnetometer, and the search coil magnetometer detects negligible AC fields from the spacecraft (even before the magnetometer booms were deployed).