PDS_VERSION_ID = PDS3 RECORD_TYPE = STREAM OBJECT = INSTRUMENT_HOST INSTRUMENT_HOST_ID = "SDU" OBJECT = INSTRUMENT_HOST_INFORMATION INSTRUMENT_HOST_NAME = "STARDUST" INSTRUMENT_HOST_TYPE = "SPACECRAFT" INSTRUMENT_HOST_DESC = " STARDUST Spacecraft Description ================================ General Configuration ===================== STARDUST is a 3-axis stabilized spacecraft designed to perform its prime mission (Wild-2 encounter) at 1.9 AU from the sun and 2.6 AU from the Earth. During the cruise periods to and from this encounter, it must be able to function adequately at a maximum distance of 2.7 AU from the sun and 3.6 AU from the Earth. The spacecraft is equipped with a power subsystem, fixed solar panels and one rechargeable battery, capable of delivering a minimum of 170 watts (W) during standard cruise operations at aphelion and a minimum of 300 W at comet encounter. The solar panels have a maximum off-sun pointing constraint of 60 degrees to avoid problems caused by refraction of light. Communications are achieved via either a high-gain, medium-gain or one of three low gain antennas. During the mission, Deep Space Network (DSN) support will be provided with primarily 34-m antennas with 70-m support being used during the close comet encounter. These antennas provide the capability for a minimum of 4000 bits per second (bps), 7900 bps expected, at encounter via the high-gain antenna and a 70-m DSN station and 40 bps at maximum Earth range via the medium gain antenna and a 34-m DSN station. The low-gain antennas, in conjunction with a 34-m DSN antenna, are ideal for near-Earth phases (Launch, Earth flyby and Earth return) when Sun-Earth-spacecraft angles are near 90 degrees, especially since they can support communications within 0.05 AU (+3 dB margin) of the Earth at a minimum data rate of 40 bps. Attitude control and propulsive maneuvers are performed using a redundant helium-fed mono-propellant (hydrazine) propulsion subsystem. The subsystem is comprised of one titanium propellant tank and a total of 16 thrusters (two strings of 8), all mounted on the lower deck of the spacecraft (opposite the high-gain antenna and solar panels - pointing toward the -z-axis of the spacecraft). Eight of these are 0.2 lb-f thrusters and are used primarily for attitude control. The other eight are 1.0 lb-f thrusters and are used for propulsive maneuvers. To avoid potential contamination of the aerogel collector, placement of thrusters on the upper deck (+z) is avoided. This configuration, however, generates uncoupled thrusts during attitude control burns and adds complexity to trajectory simulations. The normal spacecraft attitude during the mission points the +z-axis of the spacecraft to the sun. Deviations from the normal attitude are performed during communication periods and delta-velocity (DV) burns. Off-sun pointing is also permitted during non-primary science experiments, CIDA and Interstellar Dust Particle (ISP) collection, as long as the power generated by the solar arrays is adequate at the desired off-sun angle. During the comet encounter period, the +x-axis is pointed to the dust stream. These shields are placed on the spacecraft to protect it from high velocity dust impacts during comet encounter. The barriers are designed to stop a 1 cm size particle traveling at 6 km/s (which is essentially equal to the comet encounter relative velocity). Science objectives are met using three science subsystems: Aerogel Dust Collector and Sample Return Capsule (SRC), Cometary and Interstellar Dust Analyzer (CIDA) and the Dust Flux Monitor Instrument (DFMI). The imaging camera is also used for science purposes but its main function is to perform optical navigation prior to encounter with comet Wild-2. The current best estimate of the mass breakdown of the flight system is summarized in Table 1. It should be noted that this table is provided for illustrative purposes and is subject to change as per updates to the Mass Element List (MEL). Table 1 is based on Revision Z. Table 1 STARDUST Mass Element List (Rev. Z) Component Mass (kg) Component Mass (kg) S/C Power 33.378 Navigation Camera 12.686 S/C Harness 20.971 DFMI 1.530 S/C Telecom 19.222 CIDA 10.966 S/C ACS 9.951 SRC Avionics 1.992 S/C C&DH 10.394 SRC Harness 0.869 S/C Thermal 10.060 SRC Thermal 13.683 S/C Structures 104.412 SRC Structures 9.271 S/C Mechanisms 6.131 SRC Mechanisms 17.184 S/C Propulsion 19.538 SRC Parachute 4.194 Pressurant (He) 0.202 Total Dry 305.397 Propellant 85.000 Total Wet 390.599 Spacecraft Pyrotechnic Devices =============================== The Stardust spacecraft would use a total of seven (four primary, three backup) pyrotechnic devices, all classified as self-contained in case of inadvertent firing. All would be NASA Standard Initiators (NSIs) or equivalent, to be used in the deployment and release of the SRC parachute. The design includes 11 non-pyrotechnic burnwire separation devices, which would be used for releasing the SRC and solar arrays. The SRC would contain a small mortar and redundant NSIs for parachute deployment. Ordnance installation would be performed at LMA in Denver prior to shipping the spacecraft to Kennedy Space Center (KSC) for spacecraft processing. Science and Engineering Instrumentation ======================================= The scientific and engineering instruments onboard Stardust is used to gather data required to meet the science objectives. Aerogel Collector ================= Intact capture of comet dust is the highest priority of the mission. The technology in this field that has been developed over the past decade allows hypervelocity particles to be captured upon impact into an underdense, microporous media, such as silica aerogel glass. The density of the aerogel is expected to be in the range from one milligram per milliliter (mg/ml) to twenty mg/ml. The collector would have two distinct sides; the front side of the collector would be about 3 cm (1.2 in) thick, and would be used to collect the coma dust particles during flythrough. The back side that would collect the IDP would be about 1 cm (0.4 in) thick. The aerogel media would be a refinement of a series of Shuttle sample return experiments, which show that when particles with speeds greater than 3.048 km/s (10,000 ft/s, known as hypervelocity) are captured in this ultra-low density glass, they produce narrow 'carrot'-shaped tracks that are hollow and can easily be seen in the transparent aerogel. The carrot is largest at the point of entry and the particle is collected intact at the end of the carrot. Extensive experience has been gained from both laboratory and space flights of aerogel for collecting hypervelocity particles at speeds of 6 km/s (4 mi/s). With proper illumination, the track of a 10 um (3.9 x 10-4 in) particle can be seen with the naked eye. The captured particle is seen optically just beyond the tip of the carrot, and it can be recovered by a variety of techniques. Recovered samples would then be treated by sequential analysis techniques that have been developed for analysis of small meteoritic samples and IDPs. Cometary and Interplanetary Dust Analyzer (CIDA) ================================================ The CIDA would be provided to Stardust by the German Space Agency (DLR) and the Max-Planck-Institut fur Extraterrestrische Physik, with the cooperation of the Finnish Meteorological Institute and the French Space Agency (CNES). The scientific functions of the CIDA instrument are to: . determine the overall elemental composition . characterize the organic component . characterize the mineralogy of the dust particles . measure the mass and density of the dust particles . measure the range of elemental ratios . measure the light element abundance relative to carbonaceous chondrites . measure the light carbon and other gross isotopic peculiarities . measure the light element distributions within the coma . determine the mass loss from CHON particles . determine the existence and nature of very small particles, known as 'attodust' Stardust's flyby of Wild-2 would be closer than any previous comet mission. As such it is anticipated that data gathered by the CIDA could be used to increase scientific knowledge about the light element spatial distribution and the amount of mass loss from the CHON particles. Confirmation of the CHON component, in addition to characterization of the organic component, could provide better definition of chemical state and could identify more molecules. Mineralogy of individual dust particles could be gained from the additional negative ions and from more spectra. High quality spectra from the CIDA could confirm the presence of light isotopes of carbon in Halley's dust. The data would consist of more than 5000 mass spectra of individual micron- and submicron-sized particles. Ions generated by hypervelocity impact of dust onto the instrument would be separated in a drift tube and produce time-delineated mass spectra at the detector. Each spectrum would consist of a 72 microsecond (us) record digitized for transmission. Due to improved instrument design, the CIDA could be able to detect the existence of very small particles. Dust Flux Monitor Instrument (DFMI) =================================== The Dust Flux Monitor Instrument would be provided by the University of Chicago. The DFMI would measure the spatial and temporal variations of particle flux and mass distribution during the Stardust flyby of the comet. It would consist of a sensor unit (SU) and electronics box (EB) mounted to the Stardust spacecraft. The SU would be mounted to the Whipple shield, and the EB would be mounted internally to the spacecraft. The SU would consist of two independent polyvinylidene difluoride (PVDF) sensors (sensor 1 and sensor 2) mounted in a frame, and an acoustic sensor mounted one layer in from the main Whipple Shield bumper. The combined mass thresholds of sensor 1 and sensor 2 would provide cumulative and differential particle fluxes over the particle mass range of 10-11 to 10-4 gram (g) (particle diameter range 2 to 418 um), as well as cumulative flux for particles with mass greater than 1 x 10-4 g. The acoustic sensor would provide the capability of identifying impacts made by particles large enough to penetrate the Whipple Shield bumper. The EB would contain the analog and digital electronics, the instrument low voltage power supply and the SU and spacecraft interface connectors. The maximum diameter of any particle encountered during flyby is anticipated to be no larger than 1 cm (0.4 in) in diameter. For each particle impact, the DFMI would categorize the particle by size and the cumulative impact data would be transferred to the flight computer once a second. This monitoring data stream would be valuable scientific data on the concentration, size distribution and spatial distribution of dust in the coma. The accumulated dust flux would be correlated with the two-way Doppler tracking to calibrate data consistency. Specific large particle impacts would be correlated with spacecraft attitude control signals. The combined information would determine both the temporal and spatial distribution of dust sorted by the mass of the particles. Optical Navigation (OPNAV) ========================== Radiometric and optical navigation would be employed to navigate the Stardust spacecraft to Wild-2. Doppler tracking of the spacecraft would be implemented throughout all phases of the mission with this activity intensified during deep space maneuvers (DSMs), trajectory correction maneuvers, (TCMs), and the Wild-2 encounter. Optical navigation pictures of Wild-2 and adjacent stars would be taken by the spacecraft-mounted camera and would begin approximately 90 days prior to Wild-2 encounter; the rate of optical data would increase as the spacecraft approaches the comet. All navigation data would be sent to the ground for processing and maneuver command generation. The navigation facilities to be used are located at JPL. They are largely inherited from Voyager and other deep space missions carried out at JPL, and therefore no new facilities would be required. The optical navigation of the spacecraft would use the on-board star cameras as backup to acquire images of the comet and adjacent stars should the navigation camera become unavailable for this purpose. The baseline star camera would be an instrument with a 25-degree field-of-view and a 512 x 512 pixel array. The spacecraft could be navigated successfully with this instrument for the primary mission objective of cometary dust collection. Radio Frequency Lock Link ========================= From the analysis of the spacecraft telecommunications downlink signal, it is expected to be possible to obtain Doppler shifts as the spacecraft moves through the coma. This would provide a measurement of momentum transfer due to collisions with dust, and hence the spatial mass density of dust on approaching and departing from the comet nucleus. For a close flyby, it may be possible to place an upper bound on the mass of the comet nucleus from trajectory deflection. This dynamics science would utilize only that hardware needed for spacecraft telecommunications. It would not be necessary to fly any special hardware, such as an ultra-stable oscillator, to accomplish these measurement objectives. In addition, the inertial measurements unit (IMU) and star cameras could provide data on attitude changes which result from particle impacts. Analysis of this information could be useful in diagnosing large particle impact events. SAMPLE RETURN CAPSULE (SRC) =========================== Separation of the SRC from the spacecraft would occur approximately four hours prior to capsule entry. The SRC would not enter into Earth orbit, but would directly enter Earth's atmosphere, with an entry velocity of 12.8 km/s (7.9 mi/s). Taking into account SRC separation and entry corridor uncertainties, vehicle aerodynamics uncertainties and atmospheric dispersions, the landing footprint ellipse for the SRC has been determined to be approximately 84 km long by 30 km wide (52 mi x 19 mi) (three standard deviations in each direction) . [NASA 1997-A] The flight path of the SRC as it approaches UTTR would be approximately a west-northwest to east-southeast trajectory. The parachute system would consist of a mortar-deployed drogue chute to provide stability at supersonic speeds, and a main chute (8.2-m, 27-ft diameter), which would be released at about 3 km (10,000 ft). Velocity of the SRC at touchdown would be approximately 4.5 m/s (14.8 ft/s). Weight-wise, this can be compared to a 110-pound parachuter landing. At no time during the entry and descent would the SRC release its heatshield or back shell. Time elapsed from entry to touchdown would be approximately twelve minutes. Following touchdown in the early morning hours, the SRC would be recovered and transported to a staging area at UTTR in preparation for transport by NASA to the planetary materials curatorial facility at Johnson Space Center. Given the small size and mass of the SRC (see section 2.1.4.1), it is not expected that recovery and transportation of the capsule would require extraordinary handling measures or hardware other than a specialized handling fixture to be provided by LMA to cradle the capsule during transport. Other than the parachute deployment/separation system, the SRC would not contain any explosive ordnance, pyros, rocket motors, etc. Recovery Vehicle Description ============================ The SRC would be composed of five major components: heat shield, back shell, sample canister, parachute system, and avionics. The total mass of the SRC, including parachute system would be 42.6 kg (93.7 lb). The SRC would have a diameter of 81 cm (32 in). Heat Shield =========== The heat shield would be made of a graphite/epoxy composite covered with a thermal protection system (TPS). The TPS to be used for Stardust would be a phenolic impregnated carbon ablator (PICA) developed by NASA's Ames Research Center for use on high-speed reentry vehicles. The SRC heat shield would remain attached to the capsule throughout descent and serve as a protective cover for the sample canister at touchdown. Back Shell =========== The back shell structure would also be made of a graphite/epoxy composite covered with a TPS. The TPS that is planned for use on the backshell is a cork based material called SLA-561V that was developed by Lockheed Martin for use on the Viking missions to Mars and is currently used on the Space Shuttle External Tank. The back shell would provide the attach points for the parachute system. Sample Canister =============== The sample canister would be an aluminum enclosure that holds the cometary particle capture medium (aerogel) and the deployment mechanism used to deploy and stow the aerogel trays during the mission. The canister would be mounted to an equipment deck suspended between the backshell and heat shield. Parachute System ================ The parachute system would incorporate a drogue and main parachute into a single parachute canister which would contain the NSIs and the drogue deployment mortar. Inside the canister would be a gas cartridge that would be used to pressurize the mortar tube and expel the drogue chute. The drogue chute would be deployed at an altitude of approximately 30 km (100,000 ft) mean sea level (MSL) at a speed of about Mach 1.4 to provide SRC stability until the main chute is released. A gravity-switch/timer would initiate release of the drogue chute. Based on timer and backup pressure transducers, a NSI-fired cutter would release the drogue chute from the SRC at approximately 3 km (10,000 ft) MSL. As the drogue chute moves away from the SRC, it would extract the main chute from the parachute canister. Upon touchdown, cutters would fire to cut the main chute cables so that winds would not drag the SRC across the terrain. SRC Avionics ============ The current Stardust SRC baseline design includes a ultra high frequency (UHF) Locator Beacon to be used in conjunction with UHF-DF equipment on the recovery helicopters. The beacon would be turned on at main parachute deployment and would remain on until turned off by recovery personnel. The beacon would be powered by redundant sets of primary cell lithium sulfur dioxide batteries, which have long shelf life, tolerance to wide temperature extremes, and handling safety. The SRC would carry sufficient battery capacity on-board to operate the UHF beacon for at least 40 hours. Landing Footprint Determination Accuracy ======================================== The driving requirement to meet the proposed footprint onto UTTR is the entry flight path angle. An accuracy of 0.08 degrees (3 standard deviations or 3 sigma) is necessary to maintain the downrange footprint within UTTR. In order to support the trajectory accuracy needed, navigation tracking requirements have been established by JPL. Sixty-seven days prior to entry, navigation cutoff would occur and JPL would then begin to incorporate available data into a navigation solution with a highly accurate set of entry conditions. The resulting maneuver solution would be translated by LMA into a spacecraft command to be uplinked through the NASA's Deep Space Network (DSN). At entry minus sixty days the spacecraft would perform the TCM to implement preliminary entry targeting. Another TCM to refine entry targeting would be performed thirteen days prior to entry. The final entry targeting TCM would be performed 24-hours prior to entry. The landing footprint location for the Stardust capsule would be predicted by tracking the spacecraft with DSN prior to SRC release from the spacecraft. Roughly six hours prior to entry, an updated footprint would be provided to the SRC recovery management team for review of the predetermined safe entry decision criteria, which would include a one-in-a-million probability of impacting a person or damaging a range asset. Since the SRC would not have a propulsion system, there would be no way to abort the entry sequence following SRC release. Since the Stardust spacecraft would be following a direct entry path to Earth from deep space, there is only one chance for sample recovery. A no-go decision for Stardust SRC separation represents mission failure and would be considered if personnel are in danger or serious property damage is probable. Assuming the criteria for safe entry are satisfied, the SRC would be released from the spacecraft four hours before entry. The spacecraft would perform a divert maneuver at three hours prior to entry. In the event that the criteria are not met, a command would be sent to prevent SRC release, so that when the divert maneuver is performed, the entire flight system, with the SRC still attached, would be diverted from entry. Following SRC release, entry into Earth's atmosphere is defined to occur when the SRC is 125 km (77.5 mi) above a 6378 km (3,954 mi) spherical radius reference. Earth's atmosphere would quickly decelerate the SRC. The SRC trajectory would remain above 30 km (100,000 ft) until the SRC is over UTTR. Atmospheric data from high altitude weather balloons would be used for final updates to the predicted landing location." 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