- Email: [email protected]
Space technology three: mission overview and spacecraft concept description
Available online at www.sciencedirect.com
Acta Astronautica 52 (2003) 455 – 465 www.elsevier.com/locate/actaastro
Space technology three: mission overview and spacecraft concept description W.D. Deiningera;∗ , M.C. Noeckera , D.J. Wiemera , J.S. Eternoa , C. Clevenb , K.C. Patelb , G.H. Blackwoodb , L.L. Livesayb b Jet
a Ball Aerospace and Technologies Corp., P.O. Box 1062 Boulder CO 80306-1062, USA Propulsion Laboratory, California Institute of Technology, 4800 Oak Grove Drive, Pasadena CA 91109, USA
Abstract Space Technology Three (ST3), the third NASA New Millennium Program space technology mission, is focused on validation of key technologies for future separated spacecraft interferometers, such as Terrestrial Planet Finder and the Micro-Arcsecond X-ray Imaging Mission. The main technologies to be demonstrated are autonomous formation 8ying (AFF) and separated spacecraft interferometry (SSI). This mission uses two spacecraft called the Combiner and Collector, launched together, to validate these technologies. After on-orbit checkout, the spacecraft separate and are maintained within a range of 50 –1000 m. The AFF sensors and control capabilities are validated over a 2 month period followed by three months of SSI operations. Launch into an Earth-trailing heliocentric orbit is tentatively planned for March 2005 on a Delta II 7325. JPL is responsible for formation 8ying sensors and the interferometer instruments. Ball, selected as ST3 prime and working with JPL, provides the two spacecraft, system integration and test, and operations over the six-month mission duration. This paper summarizes the baseline mission concept and describes the spacecraft bus design in more detail, as de=ned at the start of the program in November 1999. c 2002 Elsevier Science Ltd. All rights reserved.
1. Introduction Human interest in =nding extrasolar planets is very strong, particularly Earth-like planets. Several missions are proposed to search for extrasolar, Earth-like planets including Terrestrial Planet Finder (TPF) [1], Terrestrial Planet Imager (TPI), and the Micro-Arcsecond X-ray Imaging Mission (MAXIM). TPF is planned for launch in 2011 and will for the =rst time isolate and analyze the light coming from the remote planet itself. The planet detection goals of TPF require a telescope arrangement some 50 –100 m ∗
Corresponding author.
long though a capability to expand the array to 100’s of meters in size would greatly enhance mission capabilities. TPI is the follow-on to TPF and will create images of the planet’s surface features. TPI might require a telescope array spanning thousands of kilometers, surely requiring a separated spacecraft approach. Each would need nanometer resolution optical path control. These missions rely on a capability to image the exo-planets directly, a diCcult activity due to the bright light emitted from the parent star. Researchers have proposed and re=ned a nulling interferometric method of selectively removing the bright starlight before detection, by superimposing the on-axis light from two telescopes (apertures) so that the stellar wavefronts interfere destructively
c 2002 Elsevier Science Ltd. All rights reserved. 0094-5765/03/$ - see front matter PII: S 0 0 9 4 - 5 7 6 5 ( 0 2 ) 0 0 1 8 8 - 1
456
W.D. Deininger et al. / Acta Astronautica 52 (2003) 455 – 465
enabling detection of the weaker oG-axis light from exo-planets [2–4]. MAXIM and its precursor, called MAXIM Path=nder, would require 2 or more spacecraft or clusters of spacecraft separated by tens to thousands of kilometers. These missions would need subnanometer precision optical path control. The New Millennium Program ST3 mission is focussed on validating separated spacecraft interferometer methods in space using two free-8ying spacecraft as the two apertures [4–7]. ST3 will launch in March 2005 into an Earth-trailing heliocentric orbit. The technologies of formation 8ying and separated spacecraft interferometry (SSI) will be demonstrated. Precision formation 8ying [8] is required to maintain proper alignments and spacings between the two spacecraft to enable use of the interferometer payload. ST3 will validate an ability to control two independent spacecraft so that their relative velocity is less than 10 microns/s while maintaining separation distances of up to 1000 m to a precision of 10 cm. Once this capability is established, ST3 will measure fringe visibility amplitudes within the interferometer as an SSI technology demonstration mission. Ball Aerospace and Technologies Corp. was selected by JPL as the prime contractor for ST3 in November 1999. JPL is responsible for the formation 8ying sensors and interferometer instruments. Ball provides the two spacecraft, system integration and test, and operations over the six month mission duration. This paper summarizes the baseline mission concept and describes the two spacecraft, as de=ned at the start of the program in November 1999, in more detail. 2. Launch and mission design Both ST3 spacecraft are stacked on a Delta II 7325 and launched. The Delta II third-stage solid motor directly injects the two spacecraft into an Earth-trailing heliocentric orbit and then separates from the two spacecraft. The spacecraft are designed and stacked to allow complete bus and payload checkout in a cluster mode. This helps to ensure a safe formation acquisition immediately after separation. After formation acquisition, the spacecraft deploy their sun-shades, and begin the month-long formation 8ying check-out and experiment period. This is
followed by the 1.5-month Combiner-mode interferometry check-out and observation period in which only the Combiner instrument is used. The baseline mission is completed by executing a three-month separated spacecraft-mode interferometer check-out and observation period. Due to the baseline dimensions of the Combiner spacecraft payload, the 3-m (10-ft) payload fairing is baselined for the ST3 mission. The mission will use the standard 3712C Delta payload attach =tting (PAF) interface. The constellation complies with center-of-gravity and thermal launch constraints for Delta II. The ST3-to-Delta electrical interface is based on the GFO design. 3. Flight system description The two-spacecraft ST3 concept is shown stacked in the launch fairing and in its operational con=guration in Fig. 1. The Combiner spacecraft has the major interferometry instrumentation on top while the Collector spacecraft shows a smaller side-mounted siderostat system. The mission drivers included formation 8ying accuracies of 10 cm over separation distances of up to 1000 m, packaging of the two spacecraft onto a single launcher with adequate mass margins, and a low onboard jitter environment to enable the interferometer to lock on and track fringes. The main performance characteristics of the two spacecraft include a total launch mass of approximately 590 kg, an attitude control capability of 0:67 arcmin, relative velocity control in formation 8ying mode of 42 m=s, jitter characteristics above 10 Hz of better than 0:05 arcsec, a power margin of 25%, and a downlink margin of 20%. The general spacecraft mass budget is shown in Table 1. Component and subsystem commonality are maximized between the two spacecraft buses to simplify integration and test and ensure low cost. Both spacecraft have identical primary structures, GNC, and spacecraft control computers. The power and cold gas propulsion systems are virtually identical, with the Combiner having a larger solar array area and two propellant tanks instead of one. Both spacecraft employ in8atable sunshades sized to reduce glint between the two spacecraft and match their mass/area ratios to within better than 1%, minimizing eGects of
W.D. Deininger et al. / Acta Astronautica 52 (2003) 455 – 465
457
Fig. 1. ST3 launch stack and operational con=guration.
Table 1 ST3 8ight system mass budget
Item
Mass (kg)
Combiner bus subsystem (dry) Combiner AFF package Combiner interferometer and electronics Combiner cold gas propellant Combiner spacecraft total (wet)
199 6 120 27 352
Collector bus subsystem (dry) Collector AFF package Collector interferometer and electronics Collector cold gas propellant Collector spacecraft total (wet)
173 6 40 15 234
Payload attach =tting Total ST3 8ight system stack mass
3 589
solar radiation pressure. Communication with Earth is through DSN using the Combiner high-gain X-band system. 3.1. General description and functional architecture The ST3 constellation consists of two three-axis stabilized spacecraft buses with integrated interferometer and autonomous formation 8ying (AFF)
sensor instrument suites. Each ST3 spacecraft consists of separate, functionally distinct subsystems— structure; mechanisms; electric power subsystem (EPS); telecommunications (TT& C); GNC; thermal control system (TCS); and propulsion—along with the payload instruments suite. The software and external interfaces make up the =nal two subsystems. Fig. 2 presents the overall ST3 Combiner spacecraft functional block diagram. The Collector spacecraft is identical except for the following diGerences: one propellant tank, 2:1 m2 of solar array area, no solid-state recorder (SSR), inclusion of the structural interface to the launcher PAF, no X-band telecommunications system, an S-band telecommunications system for emergency ground contacts only, and the Collector payload instruments. The architecture for both the Combiner and Collector spacecraft is based on Ball’s ASPEN avionics suite. ASPEN is a modular, VME-based system that uses circuit cards in a central electronics unit for all spacecraft control functions. Hardware interfaces are standardized (1553B), which simpli=es mission-speci=c modi=cation. The largely single-string ST3 con=guration provides critical redundancies resulting in high, individual spacecraft bus reliability (¿ 88%) for the 6-month mission. Hardware design is also modular, enabling parallel manufacturing and test 8ows.
458
W.D. Deininger et al. / Acta Astronautica 52 (2003) 455 – 465
Fig. 2. ST3 combiner spacecraft general block diagram.
3.2. Instrument accommodation Each spacecraft is designed to fully accommodate its ST3 instrument suite. A composite top deck provides the support needed for Combiner instruments, the required dimensional stability and stiGness, along with completely unobstructed =elds-of-view (FOV). The Collector spacecraft provides a side-mount accommodation of the Collector instrument with completely unobstructed FOV. The AFF antennas on both spacecraft are located at the edge of the solar panels to maximize horizontal and vertical separation distances without any deployments and to minimize multipath interference. The in8atable sun shade provides complete blockage of the sun over a ±45◦ angle for the inter-
ferometer instruments on both spacecraft. It has a knife-edge to minimize stray light and is made from an RF-transparent, non-re8ecting material to minimize multipath eGects. It also cants down from the mounting plane to further reduce multipath eGects and provide cleaner FOVs. Finally, the Collector sunshade is oversized to enable the mass/area ratios of the two spacecraft to be within 0.1% at each other at launch. The spacecraft bus provides all required power (including cabling and connectors), pointing, thermal, structural, and data interfaces as needed by the ST3 instrument suites and furnishes a non-contaminating environment that is critical for the sensitive optics. The use of composites enables tailoring of the interferometer mounting environment to minimize
W.D. Deininger et al. / Acta Astronautica 52 (2003) 455 – 465
459
vibrational- and thermally-induced distortions. The easily accessible external mounting interfaces for the optical instruments on both ST3 spacecraft and AFF antennas enable rapid and simple alignments during I&T. 4. Technology payloads The technology payloads consist of the interferometer and formation 8ying package (FFP). 4.1. Interferometer The interferometer instrument is described in detail elsewhere [4–7] and summarized here. The optical layout is shown in Fig. 3. The Combiner spacecraft includes a =xed-baseline, stand-alone interferometer. Here, a pair of outboard siderostats feed into an afocal gregorian compressor with a =eldstop at the internal focus. The incoming beams are then compressed and fed through delay lines, one =xed and one moveable, to the Combiner. The =xed delay line in the right arm provides a 20 m delay to compensate for the extra path length in the dogleg path of the Collector instrument. An outer annular portion of each beam is stripped oG for guiding, and the central portion of one interfering output beam used for fringe tracking. The central portion of the other beam is dispersed in a prism and integrated coherently on a CCD fringe spectrometer. The moveable delay line can be fed directly with starlight or be fed by a beam from the Collector spacecraft to accomplish SSI. In order for the Combiner/Collector interferometer system to work properly, it must have the proper on-orbit geometry. During SSI, the Combiner spacecraft resides at the focus of a parabola while the Collector spacecraft moves along the arc of the parabola (see Fig. 3). As long as the two spacecraft maintain this geometry (and the proper orientation relative to each other), the 20 m =xed delay line in the Combiner spacecraft can compensate for the dogleg path length in the Collector instrument.
Fig. 3. Optical interferometer layout and on-orbit geometry.
the delay lines can allow SSI. The FFP must measure inter-spacecraft relative distances and angles to ±1 cm and ±0:3 mrad (both to 1). Further, actuators must have small enough control authority to keep relative velocity between the two spacecraft to ¡ 10 m=s. The AFF sensor uses GPS-like signaling among multichannel transceivers on the two spacecraft [8]. Each spacecraft transmits a carrier and pseudorange signal which is received by multiple antennas on the other spacecraft. Multiple patch antennas on each spacecraft allow for both 4 steradion coverage as well as determination of relative angle and range. Fig. 3 also shows the operational geometry.
4.2. Formation 7ying package
5. Spacecraft bus subsystem designs
The FFP enables the two spacecraft to achieve a precise enough relative position and orientation that
Each of the main subsystems is described in more detail below.
460
W.D. Deininger et al. / Acta Astronautica 52 (2003) 455 – 465
5.1. Structure Both ST3 spacecraft use nearly identical, GFO-type, unibody, composite structures to minimize mass and closely match interferometer optical bench characteristics. The common primary structure is a simple, rectangular-box-shaped unibody with a single removable side panel, and top and bottom decks. The Collector spacecraft also includes a thrust tube/launch vehicle adapter cylinder that transitions the loads from the square bus cross-section to the round v-band separation system on the Delta upper stage launcher interface ring. The top deck of the Combiner spacecraft supports the interferometer payload. Radiator =ns are attached at the corners on the spacecraft to allow the removal of excess heat. Most of the spacecraft bus components are mounted on the interior and exterior surfaces of the sidewalls and removable panel. The structure is constructed of graphite/cyanate-ester composite face sheets bonded to a honeycomb core with embedded metal attach =ttings. The =xed solar panels attach to the bottom deck of each spacecraft and are stiGened with struts as shown in Fig. 1. The solar panel substrates are also made of graphite/ cyanate-ester composite facesheets and honeycomb construction as the main structure. The instruments will be mounted on the spacecraft using kinematic mounts to prevent mounting stresses on the instrument optical benches. The baseline is to have a true three-point kinematic mount consisting of mono-ball, double anti-rotation link, and single anti-rotation link. 5.2. Mechanisms During launch, the two spacecraft are held together with four electronically activated separation nuts. The separation nut uses redundant-shaped memory alloy triggers to open a segmented nut and release a bolt with up to a 2500-lb preload. After release, there are four separate spring canisters, located adjacent to the separation nuts, which provide the force to separate the spacecraft. The sunshields of both spacecraft are in8atable toruses deployed after the spacecraft separate from each other. The sunshields are attached to the outer perimeter of the solar panel substrate. Prior to deployment, they are stowed in a small toroidal
volume along the backside edge of the solar panels. The sunshields consist of thin =lm, in8atable and rigidizable materials that are deployed through the in8ation of struts using an inert gas (∼ 3 psig). The baseline rigidization technique is UV curable composite laminates due to multipath concerns. UV curable laminates are composite pre-pregs of epoxy and =berglass that are encapsulated in thin =lms and cured only by exposure to the speci=ed UV wavelength. The in8atable elements would then be vented through “T” =ttings. The gas canister and in8ation system are also attached to the backside of the solar array substrate. 5.3. Telecommunications subsystem There are two telecommunications subsystems on each ST3 spacecraft: an ultra high frequency (UHF) intersatellite link and a satellite-ground subsystem. The Combiner spacecraft has an X-band satellite-ground link to facilitate high data rate ST3 constellation earth-ground communications and is based on the small deep space transponder (SDST) while the Collector spacecraft has an S-band satellite-ground system for emergency communications only, at low data rates through the 70-m DSN. An existing UHF subsystem is selected for the intersatellite links on both spacecraft. The UHF system is sized to enable communications out to 200 km in case anomalous behavior causes the spacecraft to drift beyond the 2-km speci=cation. A detailed command and telemetry link analysis shows substantial margins for all modes of operation, as summarized in Table 2. The Combiner spacecraft RF system supports uplink rates of 250 or 7:8125 bps, along with downlink rates of 37200 or 16 bps at 0:1 AU using the 34-m DSN ground stations. The intersatellite link uses UHF transceivers and provides a data rate of 250; 000 bps, full-duplex, at 2000 m and communications capability out to 200 km in any attitude. The Collector spacecraft RF system includes a duplicate UHF intersatellite link to communicate with the Combiner. The Collector spacecraft also has the capability to communicate directly with Earth during safe mode operations using the S-band. Uplink rates from earth of 250 or 7:8125 bps and downlink rates of 16 or 7:8125 bps are supported in any spacecraft attitude at 0:1 AU using the 70-m DSN ground stations.
W.D. Deininger et al. / Acta Astronautica 52 (2003) 455 – 465
461
Table 2 ST3 telecomm links summary
Link
Link type
Data rate (bps)
Margin (dB)
Combiner uplink
Nominal—LGA 96% coverage Nominal—HGA Nominal—LGA 96% coverage Nominal—LGA 96% coverage Nominal—LGA 96% coverage Nominal worst-case
250 7.81 37,200 16 8 250 8 250 8 200,000 8
14.90 14.95 3.60 4.75 2.07 14.56 14.56 4.59 4.54 29.45 3.42
Combiner downlink Collector uplink (emergency) Collector downlink (emergency) Intersatellite link
5.4. Command and data handling (C&DH) The 1553B bus-based ASPEN C&DH architecture provides a signi=cant level of software-based, higher order processing within a modular, fully integrated spacecraft control unit (SCU) using a RAD 6000 processor. ASPEN enables all subsystems to share a central 1553B data network for system communications. The SCU localizes all processing needs in a central computer with remote data acquisition electronics. This concept diGers from previous avionics architectures by allowing higher order “processing” functionality previously accomplished in hardware to be transferred to software applications on the central computer. This transfer allowed a reduction in the number of recurring hardware components and provides functional 8exibility through data table con=gurable software applications. Each spacecraft will have an SCU which consists of a RAD 6000-based central processor unit (CPU) module, telemetry collection module (TCM), command and telemetry module (CTM), as well as modules for battery charge control (CCM), power distribution (PDM), thermal management (TM), and a low-voltage power supply (PS). Throughput capacity of the RAD 6000-based SCU is 35 MIPS. The SCU can be reprogrammed on-orbit through software updates. Interface boards in the SCU and remote interface units provide control of the entire spacecraft. The SSR, used for interim payload data storage, will be located on the Combiner spacecraft.
Critical functionality, including limited core command and telemetry services, emergency mode power control, attitude determination and control, and thermal control, is provided independent of the main processor to ensure spacecraft survival modes are not compromised. 5.5. Software ST3 will use the ASPEN 8ight software as the baseline which has been derived from the BCP2000 series 8ight software. The ASPEN 8ight software was developed using the C++ programming language for its strong exception handling capabilities. The ASPEN hardware improvements, although signi=cant, required only minimal changes to the 8ight software. The ASPEN 8ight software is supported by a catalog of software documentation, providing a detailed accounting of the full software development process. The areas where modi=cations are expected to be required include: • AFF control • Combiner-to-Collector command and telemetry links • GNC pointing and fault protection • Interferometer control capabilities • Solid-state recording capabilities 5.6. Fault protection The 8ight software protects the ST3 Combiner and Collector spacecraft against fault conditions.
462
W.D. Deininger et al. / Acta Astronautica 52 (2003) 455 – 465
Fault conditions autonomously cause fault recovery steps to be initiated, allowing the software to gracefully recover from software and/or hardware anomalous conditions. Spacecraft fault conditions such as under-voltage conditions, data transients, and single event upsets (SEU) are all supported. GNC fault conditions, such as orbit propagation errors, orbit velocity errors, attitude control calculation errors, and wheel over-speed errors are also supported. A watchdog timer is used to protect against losing processor control resulting from SEUs, operator errors, and other anomalies. The watchdog timer is continuously monitoring the 8ight software for correct operation, and if a problem occurs, it will force a processor reset and bring the spacecraft into a known safe operational state. 5.7. Guidance, navigation and control subsystem (GNC) The GNC subsystem provides continuous sixdegree-of-freedom (6-DOF) control capability during all phases of the ST3 mission. The architecture of both the Combiner and Collector spacecraft GNC are identical to provide 8exibility and reduce costs. Attitude determination is accomplished using coarse sun sensors, star trackers, a =ber-optic rate sensor, and inputs from the AFF sensor. A single Ball CT-633 star tracker per spacecraft will be the primary attitude reference for all phases of the mission. The star tracker will be mounted as closely as possible to the instrument bench to minimize alignment errors. Primary attitude actuation is accomplished using three reaction wheels. Momentum is managed in the background using thrusters, based on wheel speed limits, while observing no-thrust periods for interferometer data collection. Small wheels provide low jitter operation while providing suCcient momentum capability to compensate for solar pressure torques during no-thrust periods. A comprehensive control system analysis was performed to determine GNC performance for all ST3 mission modes. A detailed simulation of the ST3 separated spacecraft interferometry mode was created to demonstrate the compliance of the proposed ST3 GNC system with the mission requirements. In addition, a translational and rotational control system was simulated to show the ability of the proposed
Table 3 Position control performance
Parameter
Position error (cm)
Velocity error (m= s)
Minimum impulse bit AFF 3 Filtered AFF data 3 Solar radiation press 3 RSS 3 Requirement Margin
±3 ±3 ±3 ±1:4 ±4:5 ±10
±6:7 ±300 ±30 ±28 ±41:6 ±100
122%
140%
ST3 GNC system to maneuver and orient the Combiner and Collector spacecraft during this mode, while satisfying all performance parameters including quiescence during the 1000 s observation windows. Table 3 illustrates the translational control budget for the Combiner and Collector spacecraft. Solar radiation pressure forces are included, along with the assumed minimum impulse bit from the thrusters. Position errors after the separated spacecraft interferometry mode’s 1000-s thrust-free observation window are used in the budget, because this is the most demanding requirement on the position control system. 5.8. Electric power subsystem (EPS) The Ball ST3 EPS concept (identical on both spacecraft except for solar array size) exploits the ASPEN avionics architecture, leading to substantial weight savings. The performance of the EPS on each spacecraft for the main operating modes is summarized in Table 4. Power is available for each spacecraft to operate at full capacity for all phases of the mission, including the cluster-mode spacecraft check-out period. Power available in checkout mode is higher than in normal mode since the solar arrays are normal to the sun instead of at up to 45◦ angle from the sun line. The EPS is a direct energy transfer system with a battery clamped bus voltage. The system uses identical components for both the Combiner and Collector spacecraft, with the exception of the solar array shape and size. The main EPS elements include the solar array, battery, relay box, SEP with remote interface units, and the power cards mounted in the ASPEN C&DH. The power cards include battery
W.D. Deininger et al. / Acta Astronautica 52 (2003) 455 – 465 Table 4 ST3 operational-mode spacecraft bus power budget
Item
Value (W)
Combiner bus power Combiner instrument electronics Total combiner spacecraft Available combiner array power (EOL)
203 144 347 446
Combiner power margin Collector bus power Collector instrument electronics Total collector spacecraft Available collector array power (EOL) Collector power margin
99 147 60 207 259 52
charge control, bus power control and distribution, and thermal system operation. The solar arrays on both spacecraft consist of non-deployable =xed panels, oriented up to ±45◦ to the Sun. The power control circuitry (PCC) provides autonomous battery charge control, undervoltage, overvoltage, overtemperature, overpressure, and battery imbalance fault protection. The PCC also controls the fault-protected relays that switch power to the essential bus (ESB), non-essential bus (NEB), and the power switched components. The NEB supplies power to the ST3 instruments. The solar arrays provide prime power to each spacecraft in all but emergency mode. The solar panels have a total projected area of 3:2 m2 for the Combiner spacecraft and 2:1 m2 for the Collector spacecraft. The concentric con=guration allows both spacecraft to be fully powered during the clustered spacecraft check-out mode with the arrays normal to the sun. Solar array switches and power bus relays are in the relay box. The Collector spacecraft array is basically a circle 180 cm in diameter, interrupted by only the launch attach ring. The Combiner spacecraft array covers an annulus from 180 cm ID to 269 cm OD. They are populated with GaAs cells similar to the type used on GFO. The direct radiating con=guration of the arrays allows them to run cool and eCciently. The battery is used during launch to provide power to the critical spacecraft components. The battery will only be used on-orbit if the spacecraft enters the emergency mode. A nickel hydrogen battery was selected
463
as the baseline energy storage system for our ST3 spacecraft. It is a 22-cell, single-pressure vessel (SPV) unit with a rated capacity of 15 A-h. 5.9. Thermal control subsystem The TCS uses passive thermal design techniques (multilayer insulation (MLI) blankets, paint and thermal tape, and localized radiators) coupled with limited-use active heaters to allow continuous operations in all sun angles and modes of operation. The composite structure is also part of the TCS. Based on our GFO experience with composite spacecraft, we will use K1100 =bers in the lay-up of the structure to aid in moving the heat dissipated in the boxes to the radiator =ns mounted at the edges of each spacecraft. There are no moving parts. The arrangement of internal equipment is also used to aid thermal control and minimize the need for supplemental heaters. 5.10. Propulsion A trade was done to select the reaction control propulsion subsystem. Options included helium cold gas, nitrogen cold gas, pulsed plasma thrusters (PPT), =eld emission electric propulsion (FEEP), and hydrogen peroxide micro-propulsion. Drivers included the small impulse bits required (eliminated mono-prop and bi-prop), the short duration of mission operations (6 months), contamination, and the tight mission cost constraints (technology maturity). We have selected nitrogen cold gas for ST3 to maximize the reliability of demonstrating separated spacecraft formation 8ying and interferometry. The spacecraft cold-gas propulsion system consists of the propellant storage tank(s), pressure regulator, twelve 4.5-mN cold gas thrusters, latch valves, =lter and service valves along with tubing, bracketry and =eld joints. The tanks are located to minimize CG shifts during on-orbit operations and enable simple balancing within the launch fairing. The thruster locations and orientations are selected to minimize the possibility of payload equipment contamination, while allowing the necessary attitude and translation maneuvers to be conducted. The nitrogen cold gas thrusters operate at a nominal speci=c impulse of 68 s.
464
W.D. Deininger et al. / Acta Astronautica 52 (2003) 455 – 465
6. Reliability and lifetime A detailed reliability prediction was performed for the ST3 spacecraft bus con=guration. The reliability prediction was derived by comparing ST3 items with proven designs and similar hardware con=gurations used on past Ball spacecraft programs, along with speci=c vendor-provided reliability values for individual boxes. The predicted reliability for the Combiner and Collector spacecraft buses are 0.88 and 0.89, respectively. 7. Spacecraft system design features for mission technologies validation 7.1. AFF sensor validation The AFF system will be integrated with the GNC hardware and software, enabling on-orbit calibration operations to measure the accuracies of the AFF sensor outputs. The key to this process will be validating that time references and range, range rate, and bearing data achieve their speci=ed performance levels. To achieve required AFF sensor performance, the AFF antennas must have transparent lines of sight at all formation operating geometries (achieved by antenna placement and choice of spacecraft materials). Measurement errors due to the spacecraft system design must be minimized. AFF operation on orbit begins with the sequence of launch and separation actions. The two spacecraft are launched into a cluster mode, then separated, and the AFF sensor operation initialized to control the formation. Upon completion of this sequence, the Formation Flying Checkout and Experiment phase will begin. Validation of AFF system performance involves all four modes of the Formation Flying Checkout and Experiment phase, including Formation Initialization, Formation Rotation, Formation Retargeting Slew, and Formation Resizing.
Collector beyond a minimum range from the Combiner. The Collector’s thruster =rings will establish and maintain adequate range, and its reaction wheels will constrain it to attitudes suitable to avoid glint problems. 7.3. Separated spacecraft interferometry ST3 separated spacecraft interferometry requires the two spacecraft to point at a common target, and to use bearing and range data to establish the desired parabolic position oGset. Then the system must respond to =ne adjustments of position and attitude to assist in white-light fringe acquisition. These operations depend critically on AFF sensor linearity and accuracy in measuring bearing and range, which are needed to estimate optical path diGerence (OPD) and to acquire fringes quickly. The AFF system will move and retarget the Combiner and Collector spacecraft for the desired interferometer observations in the =nal three months of the mission. During the formation observation mode the AFF system will keep two spacecraft face-to-face for all observing baselines. 8. ST3 near-term plans ST3 is currently re=ning requirements and conducting architecture, instrument, and spacecraft trade studies. The project anticipates its System Architecture and Requirements Review in August of 2000. Acknowledgements The work described in this paper was developed by a team of people at Ball Aerospace & Technologies Corp. as part of our ST3 (DS3) Proposal eGort. The authors sincerely thank the Proposal Team for their work in developing this winning design.
7.2. Combiner-only interferometry The spacecraft meet the ST3 translation and angle jitter requirements, which permits the JPL instrument to perform to its requirements. In this 1.5-month phase, the AFF system must keep the
References [1] C.A. Beichman, N.J. Woolf, C.A. Lindensmith (Eds.), The Terrestrial Planet Finder (TPF)—A NASA Origins Program to Search for Habitable Planets, JPL Publication 99-3, May 1999.
W.D. Deininger et al. / Acta Astronautica 52 (2003) 455 – 465 [2] R.N. Bracewell, Detecting nonsolar planets by spinning infrared interferometer, Nature 274 (1978) 780 –781. [3] P.M. Hinz, J.R.P. Angel, W.F. HoGmann, D.W. McCarthy, Jr., P.C. McGuire, M. Cheselka, J. L. Hora, N. J. Woolf, Imaging circumstellar environments with a nulling interferometer, Nature 395 (1998) 251. [4] R.P. Lin=eld, P.W. Gorham, Science Capabilities of the ST3 Mission, Dana Point Conference, 1999. [5] K. Lau, M. Colavita, M. Shao, The new millennium separated spacecraft interferometer, Dana Point Conference, 1999.
465
[6] O.P. Lay, G.H. Blackwood, S. Dubovitsky, P.W. Gorham, R.P. Lin=eld, Design of the ST3 formation 8ying interferometer, Dana Point Conference, 1999. [7] P.W. Gorham, W.M. Folkner, G.H. Blackwood, Enabling concepts for a dual spacecraft formation-8ying optical interferometer for NASA’s ST3 Mission, Dana Point Conference, 1999. [8] G. Purcell, D. Kuang, S. Lichten, S.C. Wu, L. Young, Autonomous formation 8yer (AFF) sensor technology development, AAS 98-062, 1998.
Acta Astronautica 52 (2003) 455 – 465 www.elsevier.com/locate/actaastro
Space technology three: mission overview and spacecraft concept description W.D. Deiningera;∗ , M.C. Noeckera , D.J. Wiemera , J.S. Eternoa , C. Clevenb , K.C. Patelb , G.H. Blackwoodb , L.L. Livesayb b Jet
a Ball Aerospace and Technologies Corp., P.O. Box 1062 Boulder CO 80306-1062, USA Propulsion Laboratory, California Institute of Technology, 4800 Oak Grove Drive, Pasadena CA 91109, USA
Abstract Space Technology Three (ST3), the third NASA New Millennium Program space technology mission, is focused on validation of key technologies for future separated spacecraft interferometers, such as Terrestrial Planet Finder and the Micro-Arcsecond X-ray Imaging Mission. The main technologies to be demonstrated are autonomous formation 8ying (AFF) and separated spacecraft interferometry (SSI). This mission uses two spacecraft called the Combiner and Collector, launched together, to validate these technologies. After on-orbit checkout, the spacecraft separate and are maintained within a range of 50 –1000 m. The AFF sensors and control capabilities are validated over a 2 month period followed by three months of SSI operations. Launch into an Earth-trailing heliocentric orbit is tentatively planned for March 2005 on a Delta II 7325. JPL is responsible for formation 8ying sensors and the interferometer instruments. Ball, selected as ST3 prime and working with JPL, provides the two spacecraft, system integration and test, and operations over the six-month mission duration. This paper summarizes the baseline mission concept and describes the spacecraft bus design in more detail, as de=ned at the start of the program in November 1999. c 2002 Elsevier Science Ltd. All rights reserved.
1. Introduction Human interest in =nding extrasolar planets is very strong, particularly Earth-like planets. Several missions are proposed to search for extrasolar, Earth-like planets including Terrestrial Planet Finder (TPF) [1], Terrestrial Planet Imager (TPI), and the Micro-Arcsecond X-ray Imaging Mission (MAXIM). TPF is planned for launch in 2011 and will for the =rst time isolate and analyze the light coming from the remote planet itself. The planet detection goals of TPF require a telescope arrangement some 50 –100 m ∗
Corresponding author.
long though a capability to expand the array to 100’s of meters in size would greatly enhance mission capabilities. TPI is the follow-on to TPF and will create images of the planet’s surface features. TPI might require a telescope array spanning thousands of kilometers, surely requiring a separated spacecraft approach. Each would need nanometer resolution optical path control. These missions rely on a capability to image the exo-planets directly, a diCcult activity due to the bright light emitted from the parent star. Researchers have proposed and re=ned a nulling interferometric method of selectively removing the bright starlight before detection, by superimposing the on-axis light from two telescopes (apertures) so that the stellar wavefronts interfere destructively
c 2002 Elsevier Science Ltd. All rights reserved. 0094-5765/03/$ - see front matter PII: S 0 0 9 4 - 5 7 6 5 ( 0 2 ) 0 0 1 8 8 - 1
456
W.D. Deininger et al. / Acta Astronautica 52 (2003) 455 – 465
enabling detection of the weaker oG-axis light from exo-planets [2–4]. MAXIM and its precursor, called MAXIM Path=nder, would require 2 or more spacecraft or clusters of spacecraft separated by tens to thousands of kilometers. These missions would need subnanometer precision optical path control. The New Millennium Program ST3 mission is focussed on validating separated spacecraft interferometer methods in space using two free-8ying spacecraft as the two apertures [4–7]. ST3 will launch in March 2005 into an Earth-trailing heliocentric orbit. The technologies of formation 8ying and separated spacecraft interferometry (SSI) will be demonstrated. Precision formation 8ying [8] is required to maintain proper alignments and spacings between the two spacecraft to enable use of the interferometer payload. ST3 will validate an ability to control two independent spacecraft so that their relative velocity is less than 10 microns/s while maintaining separation distances of up to 1000 m to a precision of 10 cm. Once this capability is established, ST3 will measure fringe visibility amplitudes within the interferometer as an SSI technology demonstration mission. Ball Aerospace and Technologies Corp. was selected by JPL as the prime contractor for ST3 in November 1999. JPL is responsible for the formation 8ying sensors and interferometer instruments. Ball provides the two spacecraft, system integration and test, and operations over the six month mission duration. This paper summarizes the baseline mission concept and describes the two spacecraft, as de=ned at the start of the program in November 1999, in more detail. 2. Launch and mission design Both ST3 spacecraft are stacked on a Delta II 7325 and launched. The Delta II third-stage solid motor directly injects the two spacecraft into an Earth-trailing heliocentric orbit and then separates from the two spacecraft. The spacecraft are designed and stacked to allow complete bus and payload checkout in a cluster mode. This helps to ensure a safe formation acquisition immediately after separation. After formation acquisition, the spacecraft deploy their sun-shades, and begin the month-long formation 8ying check-out and experiment period. This is
followed by the 1.5-month Combiner-mode interferometry check-out and observation period in which only the Combiner instrument is used. The baseline mission is completed by executing a three-month separated spacecraft-mode interferometer check-out and observation period. Due to the baseline dimensions of the Combiner spacecraft payload, the 3-m (10-ft) payload fairing is baselined for the ST3 mission. The mission will use the standard 3712C Delta payload attach =tting (PAF) interface. The constellation complies with center-of-gravity and thermal launch constraints for Delta II. The ST3-to-Delta electrical interface is based on the GFO design. 3. Flight system description The two-spacecraft ST3 concept is shown stacked in the launch fairing and in its operational con=guration in Fig. 1. The Combiner spacecraft has the major interferometry instrumentation on top while the Collector spacecraft shows a smaller side-mounted siderostat system. The mission drivers included formation 8ying accuracies of 10 cm over separation distances of up to 1000 m, packaging of the two spacecraft onto a single launcher with adequate mass margins, and a low onboard jitter environment to enable the interferometer to lock on and track fringes. The main performance characteristics of the two spacecraft include a total launch mass of approximately 590 kg, an attitude control capability of 0:67 arcmin, relative velocity control in formation 8ying mode of 42 m=s, jitter characteristics above 10 Hz of better than 0:05 arcsec, a power margin of 25%, and a downlink margin of 20%. The general spacecraft mass budget is shown in Table 1. Component and subsystem commonality are maximized between the two spacecraft buses to simplify integration and test and ensure low cost. Both spacecraft have identical primary structures, GNC, and spacecraft control computers. The power and cold gas propulsion systems are virtually identical, with the Combiner having a larger solar array area and two propellant tanks instead of one. Both spacecraft employ in8atable sunshades sized to reduce glint between the two spacecraft and match their mass/area ratios to within better than 1%, minimizing eGects of
W.D. Deininger et al. / Acta Astronautica 52 (2003) 455 – 465
457
Fig. 1. ST3 launch stack and operational con=guration.
Table 1 ST3 8ight system mass budget
Item
Mass (kg)
Combiner bus subsystem (dry) Combiner AFF package Combiner interferometer and electronics Combiner cold gas propellant Combiner spacecraft total (wet)
199 6 120 27 352
Collector bus subsystem (dry) Collector AFF package Collector interferometer and electronics Collector cold gas propellant Collector spacecraft total (wet)
173 6 40 15 234
Payload attach =tting Total ST3 8ight system stack mass
3 589
solar radiation pressure. Communication with Earth is through DSN using the Combiner high-gain X-band system. 3.1. General description and functional architecture The ST3 constellation consists of two three-axis stabilized spacecraft buses with integrated interferometer and autonomous formation 8ying (AFF)
sensor instrument suites. Each ST3 spacecraft consists of separate, functionally distinct subsystems— structure; mechanisms; electric power subsystem (EPS); telecommunications (TT& C); GNC; thermal control system (TCS); and propulsion—along with the payload instruments suite. The software and external interfaces make up the =nal two subsystems. Fig. 2 presents the overall ST3 Combiner spacecraft functional block diagram. The Collector spacecraft is identical except for the following diGerences: one propellant tank, 2:1 m2 of solar array area, no solid-state recorder (SSR), inclusion of the structural interface to the launcher PAF, no X-band telecommunications system, an S-band telecommunications system for emergency ground contacts only, and the Collector payload instruments. The architecture for both the Combiner and Collector spacecraft is based on Ball’s ASPEN avionics suite. ASPEN is a modular, VME-based system that uses circuit cards in a central electronics unit for all spacecraft control functions. Hardware interfaces are standardized (1553B), which simpli=es mission-speci=c modi=cation. The largely single-string ST3 con=guration provides critical redundancies resulting in high, individual spacecraft bus reliability (¿ 88%) for the 6-month mission. Hardware design is also modular, enabling parallel manufacturing and test 8ows.
458
W.D. Deininger et al. / Acta Astronautica 52 (2003) 455 – 465
Fig. 2. ST3 combiner spacecraft general block diagram.
3.2. Instrument accommodation Each spacecraft is designed to fully accommodate its ST3 instrument suite. A composite top deck provides the support needed for Combiner instruments, the required dimensional stability and stiGness, along with completely unobstructed =elds-of-view (FOV). The Collector spacecraft provides a side-mount accommodation of the Collector instrument with completely unobstructed FOV. The AFF antennas on both spacecraft are located at the edge of the solar panels to maximize horizontal and vertical separation distances without any deployments and to minimize multipath interference. The in8atable sun shade provides complete blockage of the sun over a ±45◦ angle for the inter-
ferometer instruments on both spacecraft. It has a knife-edge to minimize stray light and is made from an RF-transparent, non-re8ecting material to minimize multipath eGects. It also cants down from the mounting plane to further reduce multipath eGects and provide cleaner FOVs. Finally, the Collector sunshade is oversized to enable the mass/area ratios of the two spacecraft to be within 0.1% at each other at launch. The spacecraft bus provides all required power (including cabling and connectors), pointing, thermal, structural, and data interfaces as needed by the ST3 instrument suites and furnishes a non-contaminating environment that is critical for the sensitive optics. The use of composites enables tailoring of the interferometer mounting environment to minimize
W.D. Deininger et al. / Acta Astronautica 52 (2003) 455 – 465
459
vibrational- and thermally-induced distortions. The easily accessible external mounting interfaces for the optical instruments on both ST3 spacecraft and AFF antennas enable rapid and simple alignments during I&T. 4. Technology payloads The technology payloads consist of the interferometer and formation 8ying package (FFP). 4.1. Interferometer The interferometer instrument is described in detail elsewhere [4–7] and summarized here. The optical layout is shown in Fig. 3. The Combiner spacecraft includes a =xed-baseline, stand-alone interferometer. Here, a pair of outboard siderostats feed into an afocal gregorian compressor with a =eldstop at the internal focus. The incoming beams are then compressed and fed through delay lines, one =xed and one moveable, to the Combiner. The =xed delay line in the right arm provides a 20 m delay to compensate for the extra path length in the dogleg path of the Collector instrument. An outer annular portion of each beam is stripped oG for guiding, and the central portion of one interfering output beam used for fringe tracking. The central portion of the other beam is dispersed in a prism and integrated coherently on a CCD fringe spectrometer. The moveable delay line can be fed directly with starlight or be fed by a beam from the Collector spacecraft to accomplish SSI. In order for the Combiner/Collector interferometer system to work properly, it must have the proper on-orbit geometry. During SSI, the Combiner spacecraft resides at the focus of a parabola while the Collector spacecraft moves along the arc of the parabola (see Fig. 3). As long as the two spacecraft maintain this geometry (and the proper orientation relative to each other), the 20 m =xed delay line in the Combiner spacecraft can compensate for the dogleg path length in the Collector instrument.
Fig. 3. Optical interferometer layout and on-orbit geometry.
the delay lines can allow SSI. The FFP must measure inter-spacecraft relative distances and angles to ±1 cm and ±0:3 mrad (both to 1). Further, actuators must have small enough control authority to keep relative velocity between the two spacecraft to ¡ 10 m=s. The AFF sensor uses GPS-like signaling among multichannel transceivers on the two spacecraft [8]. Each spacecraft transmits a carrier and pseudorange signal which is received by multiple antennas on the other spacecraft. Multiple patch antennas on each spacecraft allow for both 4 steradion coverage as well as determination of relative angle and range. Fig. 3 also shows the operational geometry.
4.2. Formation 7ying package
5. Spacecraft bus subsystem designs
The FFP enables the two spacecraft to achieve a precise enough relative position and orientation that
Each of the main subsystems is described in more detail below.
460
W.D. Deininger et al. / Acta Astronautica 52 (2003) 455 – 465
5.1. Structure Both ST3 spacecraft use nearly identical, GFO-type, unibody, composite structures to minimize mass and closely match interferometer optical bench characteristics. The common primary structure is a simple, rectangular-box-shaped unibody with a single removable side panel, and top and bottom decks. The Collector spacecraft also includes a thrust tube/launch vehicle adapter cylinder that transitions the loads from the square bus cross-section to the round v-band separation system on the Delta upper stage launcher interface ring. The top deck of the Combiner spacecraft supports the interferometer payload. Radiator =ns are attached at the corners on the spacecraft to allow the removal of excess heat. Most of the spacecraft bus components are mounted on the interior and exterior surfaces of the sidewalls and removable panel. The structure is constructed of graphite/cyanate-ester composite face sheets bonded to a honeycomb core with embedded metal attach =ttings. The =xed solar panels attach to the bottom deck of each spacecraft and are stiGened with struts as shown in Fig. 1. The solar panel substrates are also made of graphite/ cyanate-ester composite facesheets and honeycomb construction as the main structure. The instruments will be mounted on the spacecraft using kinematic mounts to prevent mounting stresses on the instrument optical benches. The baseline is to have a true three-point kinematic mount consisting of mono-ball, double anti-rotation link, and single anti-rotation link. 5.2. Mechanisms During launch, the two spacecraft are held together with four electronically activated separation nuts. The separation nut uses redundant-shaped memory alloy triggers to open a segmented nut and release a bolt with up to a 2500-lb preload. After release, there are four separate spring canisters, located adjacent to the separation nuts, which provide the force to separate the spacecraft. The sunshields of both spacecraft are in8atable toruses deployed after the spacecraft separate from each other. The sunshields are attached to the outer perimeter of the solar panel substrate. Prior to deployment, they are stowed in a small toroidal
volume along the backside edge of the solar panels. The sunshields consist of thin =lm, in8atable and rigidizable materials that are deployed through the in8ation of struts using an inert gas (∼ 3 psig). The baseline rigidization technique is UV curable composite laminates due to multipath concerns. UV curable laminates are composite pre-pregs of epoxy and =berglass that are encapsulated in thin =lms and cured only by exposure to the speci=ed UV wavelength. The in8atable elements would then be vented through “T” =ttings. The gas canister and in8ation system are also attached to the backside of the solar array substrate. 5.3. Telecommunications subsystem There are two telecommunications subsystems on each ST3 spacecraft: an ultra high frequency (UHF) intersatellite link and a satellite-ground subsystem. The Combiner spacecraft has an X-band satellite-ground link to facilitate high data rate ST3 constellation earth-ground communications and is based on the small deep space transponder (SDST) while the Collector spacecraft has an S-band satellite-ground system for emergency communications only, at low data rates through the 70-m DSN. An existing UHF subsystem is selected for the intersatellite links on both spacecraft. The UHF system is sized to enable communications out to 200 km in case anomalous behavior causes the spacecraft to drift beyond the 2-km speci=cation. A detailed command and telemetry link analysis shows substantial margins for all modes of operation, as summarized in Table 2. The Combiner spacecraft RF system supports uplink rates of 250 or 7:8125 bps, along with downlink rates of 37200 or 16 bps at 0:1 AU using the 34-m DSN ground stations. The intersatellite link uses UHF transceivers and provides a data rate of 250; 000 bps, full-duplex, at 2000 m and communications capability out to 200 km in any attitude. The Collector spacecraft RF system includes a duplicate UHF intersatellite link to communicate with the Combiner. The Collector spacecraft also has the capability to communicate directly with Earth during safe mode operations using the S-band. Uplink rates from earth of 250 or 7:8125 bps and downlink rates of 16 or 7:8125 bps are supported in any spacecraft attitude at 0:1 AU using the 70-m DSN ground stations.
W.D. Deininger et al. / Acta Astronautica 52 (2003) 455 – 465
461
Table 2 ST3 telecomm links summary
Link
Link type
Data rate (bps)
Margin (dB)
Combiner uplink
Nominal—LGA 96% coverage Nominal—HGA Nominal—LGA 96% coverage Nominal—LGA 96% coverage Nominal—LGA 96% coverage Nominal worst-case
250 7.81 37,200 16 8 250 8 250 8 200,000 8
14.90 14.95 3.60 4.75 2.07 14.56 14.56 4.59 4.54 29.45 3.42
Combiner downlink Collector uplink (emergency) Collector downlink (emergency) Intersatellite link
5.4. Command and data handling (C&DH) The 1553B bus-based ASPEN C&DH architecture provides a signi=cant level of software-based, higher order processing within a modular, fully integrated spacecraft control unit (SCU) using a RAD 6000 processor. ASPEN enables all subsystems to share a central 1553B data network for system communications. The SCU localizes all processing needs in a central computer with remote data acquisition electronics. This concept diGers from previous avionics architectures by allowing higher order “processing” functionality previously accomplished in hardware to be transferred to software applications on the central computer. This transfer allowed a reduction in the number of recurring hardware components and provides functional 8exibility through data table con=gurable software applications. Each spacecraft will have an SCU which consists of a RAD 6000-based central processor unit (CPU) module, telemetry collection module (TCM), command and telemetry module (CTM), as well as modules for battery charge control (CCM), power distribution (PDM), thermal management (TM), and a low-voltage power supply (PS). Throughput capacity of the RAD 6000-based SCU is 35 MIPS. The SCU can be reprogrammed on-orbit through software updates. Interface boards in the SCU and remote interface units provide control of the entire spacecraft. The SSR, used for interim payload data storage, will be located on the Combiner spacecraft.
Critical functionality, including limited core command and telemetry services, emergency mode power control, attitude determination and control, and thermal control, is provided independent of the main processor to ensure spacecraft survival modes are not compromised. 5.5. Software ST3 will use the ASPEN 8ight software as the baseline which has been derived from the BCP2000 series 8ight software. The ASPEN 8ight software was developed using the C++ programming language for its strong exception handling capabilities. The ASPEN hardware improvements, although signi=cant, required only minimal changes to the 8ight software. The ASPEN 8ight software is supported by a catalog of software documentation, providing a detailed accounting of the full software development process. The areas where modi=cations are expected to be required include: • AFF control • Combiner-to-Collector command and telemetry links • GNC pointing and fault protection • Interferometer control capabilities • Solid-state recording capabilities 5.6. Fault protection The 8ight software protects the ST3 Combiner and Collector spacecraft against fault conditions.
462
W.D. Deininger et al. / Acta Astronautica 52 (2003) 455 – 465
Fault conditions autonomously cause fault recovery steps to be initiated, allowing the software to gracefully recover from software and/or hardware anomalous conditions. Spacecraft fault conditions such as under-voltage conditions, data transients, and single event upsets (SEU) are all supported. GNC fault conditions, such as orbit propagation errors, orbit velocity errors, attitude control calculation errors, and wheel over-speed errors are also supported. A watchdog timer is used to protect against losing processor control resulting from SEUs, operator errors, and other anomalies. The watchdog timer is continuously monitoring the 8ight software for correct operation, and if a problem occurs, it will force a processor reset and bring the spacecraft into a known safe operational state. 5.7. Guidance, navigation and control subsystem (GNC) The GNC subsystem provides continuous sixdegree-of-freedom (6-DOF) control capability during all phases of the ST3 mission. The architecture of both the Combiner and Collector spacecraft GNC are identical to provide 8exibility and reduce costs. Attitude determination is accomplished using coarse sun sensors, star trackers, a =ber-optic rate sensor, and inputs from the AFF sensor. A single Ball CT-633 star tracker per spacecraft will be the primary attitude reference for all phases of the mission. The star tracker will be mounted as closely as possible to the instrument bench to minimize alignment errors. Primary attitude actuation is accomplished using three reaction wheels. Momentum is managed in the background using thrusters, based on wheel speed limits, while observing no-thrust periods for interferometer data collection. Small wheels provide low jitter operation while providing suCcient momentum capability to compensate for solar pressure torques during no-thrust periods. A comprehensive control system analysis was performed to determine GNC performance for all ST3 mission modes. A detailed simulation of the ST3 separated spacecraft interferometry mode was created to demonstrate the compliance of the proposed ST3 GNC system with the mission requirements. In addition, a translational and rotational control system was simulated to show the ability of the proposed
Table 3 Position control performance
Parameter
Position error (cm)
Velocity error (m= s)
Minimum impulse bit AFF 3 Filtered AFF data 3 Solar radiation press 3 RSS 3 Requirement Margin
±3 ±3 ±3 ±1:4 ±4:5 ±10
±6:7 ±300 ±30 ±28 ±41:6 ±100
122%
140%
ST3 GNC system to maneuver and orient the Combiner and Collector spacecraft during this mode, while satisfying all performance parameters including quiescence during the 1000 s observation windows. Table 3 illustrates the translational control budget for the Combiner and Collector spacecraft. Solar radiation pressure forces are included, along with the assumed minimum impulse bit from the thrusters. Position errors after the separated spacecraft interferometry mode’s 1000-s thrust-free observation window are used in the budget, because this is the most demanding requirement on the position control system. 5.8. Electric power subsystem (EPS) The Ball ST3 EPS concept (identical on both spacecraft except for solar array size) exploits the ASPEN avionics architecture, leading to substantial weight savings. The performance of the EPS on each spacecraft for the main operating modes is summarized in Table 4. Power is available for each spacecraft to operate at full capacity for all phases of the mission, including the cluster-mode spacecraft check-out period. Power available in checkout mode is higher than in normal mode since the solar arrays are normal to the sun instead of at up to 45◦ angle from the sun line. The EPS is a direct energy transfer system with a battery clamped bus voltage. The system uses identical components for both the Combiner and Collector spacecraft, with the exception of the solar array shape and size. The main EPS elements include the solar array, battery, relay box, SEP with remote interface units, and the power cards mounted in the ASPEN C&DH. The power cards include battery
W.D. Deininger et al. / Acta Astronautica 52 (2003) 455 – 465 Table 4 ST3 operational-mode spacecraft bus power budget
Item
Value (W)
Combiner bus power Combiner instrument electronics Total combiner spacecraft Available combiner array power (EOL)
203 144 347 446
Combiner power margin Collector bus power Collector instrument electronics Total collector spacecraft Available collector array power (EOL) Collector power margin
99 147 60 207 259 52
charge control, bus power control and distribution, and thermal system operation. The solar arrays on both spacecraft consist of non-deployable =xed panels, oriented up to ±45◦ to the Sun. The power control circuitry (PCC) provides autonomous battery charge control, undervoltage, overvoltage, overtemperature, overpressure, and battery imbalance fault protection. The PCC also controls the fault-protected relays that switch power to the essential bus (ESB), non-essential bus (NEB), and the power switched components. The NEB supplies power to the ST3 instruments. The solar arrays provide prime power to each spacecraft in all but emergency mode. The solar panels have a total projected area of 3:2 m2 for the Combiner spacecraft and 2:1 m2 for the Collector spacecraft. The concentric con=guration allows both spacecraft to be fully powered during the clustered spacecraft check-out mode with the arrays normal to the sun. Solar array switches and power bus relays are in the relay box. The Collector spacecraft array is basically a circle 180 cm in diameter, interrupted by only the launch attach ring. The Combiner spacecraft array covers an annulus from 180 cm ID to 269 cm OD. They are populated with GaAs cells similar to the type used on GFO. The direct radiating con=guration of the arrays allows them to run cool and eCciently. The battery is used during launch to provide power to the critical spacecraft components. The battery will only be used on-orbit if the spacecraft enters the emergency mode. A nickel hydrogen battery was selected
463
as the baseline energy storage system for our ST3 spacecraft. It is a 22-cell, single-pressure vessel (SPV) unit with a rated capacity of 15 A-h. 5.9. Thermal control subsystem The TCS uses passive thermal design techniques (multilayer insulation (MLI) blankets, paint and thermal tape, and localized radiators) coupled with limited-use active heaters to allow continuous operations in all sun angles and modes of operation. The composite structure is also part of the TCS. Based on our GFO experience with composite spacecraft, we will use K1100 =bers in the lay-up of the structure to aid in moving the heat dissipated in the boxes to the radiator =ns mounted at the edges of each spacecraft. There are no moving parts. The arrangement of internal equipment is also used to aid thermal control and minimize the need for supplemental heaters. 5.10. Propulsion A trade was done to select the reaction control propulsion subsystem. Options included helium cold gas, nitrogen cold gas, pulsed plasma thrusters (PPT), =eld emission electric propulsion (FEEP), and hydrogen peroxide micro-propulsion. Drivers included the small impulse bits required (eliminated mono-prop and bi-prop), the short duration of mission operations (6 months), contamination, and the tight mission cost constraints (technology maturity). We have selected nitrogen cold gas for ST3 to maximize the reliability of demonstrating separated spacecraft formation 8ying and interferometry. The spacecraft cold-gas propulsion system consists of the propellant storage tank(s), pressure regulator, twelve 4.5-mN cold gas thrusters, latch valves, =lter and service valves along with tubing, bracketry and =eld joints. The tanks are located to minimize CG shifts during on-orbit operations and enable simple balancing within the launch fairing. The thruster locations and orientations are selected to minimize the possibility of payload equipment contamination, while allowing the necessary attitude and translation maneuvers to be conducted. The nitrogen cold gas thrusters operate at a nominal speci=c impulse of 68 s.
464
W.D. Deininger et al. / Acta Astronautica 52 (2003) 455 – 465
6. Reliability and lifetime A detailed reliability prediction was performed for the ST3 spacecraft bus con=guration. The reliability prediction was derived by comparing ST3 items with proven designs and similar hardware con=gurations used on past Ball spacecraft programs, along with speci=c vendor-provided reliability values for individual boxes. The predicted reliability for the Combiner and Collector spacecraft buses are 0.88 and 0.89, respectively. 7. Spacecraft system design features for mission technologies validation 7.1. AFF sensor validation The AFF system will be integrated with the GNC hardware and software, enabling on-orbit calibration operations to measure the accuracies of the AFF sensor outputs. The key to this process will be validating that time references and range, range rate, and bearing data achieve their speci=ed performance levels. To achieve required AFF sensor performance, the AFF antennas must have transparent lines of sight at all formation operating geometries (achieved by antenna placement and choice of spacecraft materials). Measurement errors due to the spacecraft system design must be minimized. AFF operation on orbit begins with the sequence of launch and separation actions. The two spacecraft are launched into a cluster mode, then separated, and the AFF sensor operation initialized to control the formation. Upon completion of this sequence, the Formation Flying Checkout and Experiment phase will begin. Validation of AFF system performance involves all four modes of the Formation Flying Checkout and Experiment phase, including Formation Initialization, Formation Rotation, Formation Retargeting Slew, and Formation Resizing.
Collector beyond a minimum range from the Combiner. The Collector’s thruster =rings will establish and maintain adequate range, and its reaction wheels will constrain it to attitudes suitable to avoid glint problems. 7.3. Separated spacecraft interferometry ST3 separated spacecraft interferometry requires the two spacecraft to point at a common target, and to use bearing and range data to establish the desired parabolic position oGset. Then the system must respond to =ne adjustments of position and attitude to assist in white-light fringe acquisition. These operations depend critically on AFF sensor linearity and accuracy in measuring bearing and range, which are needed to estimate optical path diGerence (OPD) and to acquire fringes quickly. The AFF system will move and retarget the Combiner and Collector spacecraft for the desired interferometer observations in the =nal three months of the mission. During the formation observation mode the AFF system will keep two spacecraft face-to-face for all observing baselines. 8. ST3 near-term plans ST3 is currently re=ning requirements and conducting architecture, instrument, and spacecraft trade studies. The project anticipates its System Architecture and Requirements Review in August of 2000. Acknowledgements The work described in this paper was developed by a team of people at Ball Aerospace & Technologies Corp. as part of our ST3 (DS3) Proposal eGort. The authors sincerely thank the Proposal Team for their work in developing this winning design.
7.2. Combiner-only interferometry The spacecraft meet the ST3 translation and angle jitter requirements, which permits the JPL instrument to perform to its requirements. In this 1.5-month phase, the AFF system must keep the
References [1] C.A. Beichman, N.J. Woolf, C.A. Lindensmith (Eds.), The Terrestrial Planet Finder (TPF)—A NASA Origins Program to Search for Habitable Planets, JPL Publication 99-3, May 1999.
W.D. Deininger et al. / Acta Astronautica 52 (2003) 455 – 465 [2] R.N. Bracewell, Detecting nonsolar planets by spinning infrared interferometer, Nature 274 (1978) 780 –781. [3] P.M. Hinz, J.R.P. Angel, W.F. HoGmann, D.W. McCarthy, Jr., P.C. McGuire, M. Cheselka, J. L. Hora, N. J. Woolf, Imaging circumstellar environments with a nulling interferometer, Nature 395 (1998) 251. [4] R.P. Lin=eld, P.W. Gorham, Science Capabilities of the ST3 Mission, Dana Point Conference, 1999. [5] K. Lau, M. Colavita, M. Shao, The new millennium separated spacecraft interferometer, Dana Point Conference, 1999.
465
[6] O.P. Lay, G.H. Blackwood, S. Dubovitsky, P.W. Gorham, R.P. Lin=eld, Design of the ST3 formation 8ying interferometer, Dana Point Conference, 1999. [7] P.W. Gorham, W.M. Folkner, G.H. Blackwood, Enabling concepts for a dual spacecraft formation-8ying optical interferometer for NASA’s ST3 Mission, Dana Point Conference, 1999. [8] G. Purcell, D. Kuang, S. Lichten, S.C. Wu, L. Young, Autonomous formation 8yer (AFF) sensor technology development, AAS 98-062, 1998.