Deep Space 1
DS1 (Deep Space 1)
DS1 is the first satellite mission in NASA's New Millennium Program (NMP). A major objective of NMP is to identify and to demonstrate new technologies in spacecraft and instrument design as well as in spacecraft operations, to validate new and high-risk concepts in the spaceborne environment, thereby advancing the horizons of future missions. DS1 is in particular a low-cost minisatellite technology demonstration mission of NASA/JPL with the following objectives: 1) 2) 3) 4) 5)
1) Demonstrate the in-space flight operations and quantify the performance of the following five advanced technologies:
- Solar electric propulsion (SEP)
- Solar concentrator arrays
- Autonomous navigation
- Miniature camera and imaging spectrometer
- Small deep-space transponder
In addition, the objectives called for the testing of any three of the following six advance technologies:
- Ka-band solid-state power amplifier
- Beacon monitor operations
- Autonomous remote agent
- Low-power electronics
- Power actuation and switching module
- Multifunctional Structure (MFS)
2) Acquire the data necessary to quantify the performance of these advanced technologies by the end of the primary mission (Sept. 1999).
3) Utilize the on-board IPS to propel the DS1 spacecraft on a trajectory to encounter an asteroid in 1999.
4) Assess the interaction of the IPS operations with the S/C and its potential impact on charged particle, radio waves and plasma, and other science investigations on future SEP-propelled missions. - A twelfth technology, a miniature integrated ion and electron spectrometer, PEPE, was not included in the mission success criteria, but it was on board and received a thorough evaluation.
Figure 1: Artist's view of the DS1 spacecraft validating ion engine propulsion technologies (image credit: NASA/JPL)
The structure of the minisatellite consists of an octagonal aluminum space frame (1.1 m x 1.1 m x 1.5 high), the overall stowed S/C dimensions are: 2.5 m high, 2.1 m deep, and 1.7 m wide, based on the MSTI (Miniature Seeker Technology Integration) spacecraft, built by Spectrum Astro, Inc. of Gilbert, AZ (SA-200HP bus). With most of the components mounted on the exterior of the bus, their accessibility simplifies replacement during integration and test.
The S/C is powered by batteries and two solar panels attached to the sides of the frame which span of 11.75 m when deployed. The solar panels, designated SCARLET II (Solar Concentrator Arrays with Refractive Linear Element Technology) constitute one of the technology tests on the spacecraft (see SCARLET below). The pair of solar arrays produce 2.5 kW of power at 1AU. Each array is comprised of four panels and a single-axis gimbal controls to point the panels in the more sensitive longitudinal axis toward the sun.
The total mass of the S/C is about 486.3 kg, composed of 373.7 kg dry spacecraft, 31.1 kg of hydrazine, and 81.5 kg of xenon for IPS. A MIL-STD-1553 data bus is used for onboard communications.
Figure 2: Front and back view of the DS1 spacecraft in stowed configuration (image credit: NASA/JPL)
Figure 3: Illustration of the deployed DS1 spacecraft (image credit: NASA/JPL)
RF communications are via a high gain antenna, three low gain antennas, and a Ka-band antenna, mounted on top of the spacecraft. One low-gain antenna is mounted on the service boom and points toward Earth. The CCSDS protocol suite is used for all data communications.
Launch: The launch of DS1 took place on October 24, 1998 on a Delta launcher (7326-9.5 with three strap-on solid propellant rockets) from Cape Canaveral Air Station, FL (secondary payload: SEDSAT-1).
Its first destination was the near-Earth asteroid “1992 KD” with an estimated diameter of 2-5 km. The primary mission of DS1 ended on September 29, 1999, completing its mission of demonstrating the new technologies.
Orbit: The SEDSAT-1 microsatellite (University of Alabama, Huntsville) was ejected from the second stage. The Delta vehicle entered a 185 km parking orbit, then fired again to enter an orbit: 174 km x 2744 km, inclination of 28.5º. The third stage separated and accelerated to a solar orbit with the DS1 satellite, while the second stage burned again with SEDSAT-1 for an orbit of: 556 km x 1042 km, inclination of 31.5º.
DS1 was injected into a solar orbit with the third stage burn. Orbital period: 453.00 days, inclination = 0.4º, periapsis = 0.99 AU, apoapsis = 1.32 AU, eccentricity = 0.14300, epoch start date/time = 1998.297:12:08:00 (24 Oct. 1998).
Figure 4: Orbits of DS1 and asteroid 1992 KD for rendezvous of primary mission (image credit: NASA/JPL)
Figure 5: DS1 trajectory for the primary mission (through Sept. 28, 1999) and extended mission (image credit: NASA/JPL)
Advanced technology payload complement
The DS1 payload consists of 12 technologies, two of them happen to be science instruments.
IPS (Ion Propulsion System):
Objective: validation of system performance. IPS, provided by the project NSTAR [(NASA SEP Technology Application Readiness), where SEP is the acronym for Solar Electric Propulsion], uses a hollow cathode to produce electrons to ionize xenon by collision. The NSTAR/IPS system consists of a 30 cm xenon ion thruster, xenon feed system (XFS), Power Processing Unit (PPU), and a Digital Control and Interface Unit (DCIU).
The Xe+ is electrostatically accelerated through a potential of up to 1280 V and emitted from the 30 cm diameter thruster through a pair of molybdenum grids. A separate electron beam is emitted to produce a neutral plasma beam. The power processing unit (PPU) of the IPS can accept as much as 2.5 kW, corresponding to a peak thruster operating power of 2.3 kW and a thrust of 92 mN. Throttling is achieved by balancing thruster and Xe feed system parameters at lower power levels; and at the lowest PPU thruster power, 525 W, the thrust is 19 mN. The specific impulse decreases from 31,400 m/s (or 3300 s) at high power to 19,000 m/s at the minimum throttle level.
A comprehensive diagnostic system is also on the spacecraft to validate system performance. The diagnostic instrument suite includes a retarding potential analyzer, two Langmuir probes, search-coil and fluxgate magnetometers, a plasma wave sensor, and two pairs of quartz-crystal microbalances and calorimeters.
Figure 6: Schematic illustration of the IPS elements (image credit: NASA)
At the end of the primary mission in Sept. 1999, IPS had operated for about 3000 hours. The IPS operated over a broad range of its 112 throttle levels, from input power levels of 580 W to 2140 W, with corresponding specific impulses of 19,375 m/s and 31,200 m/s, respectively (determined by radio navigation). Up to June 30, 1999, there were 1799.4 hours of thrusting by IPS, the total Xe consumption was 11.4 kg, providing 699.6 m/s. - By August 31, 2000 IPS had accumulated > 5200 hours of operations, a longer time period than any propulsion system on any S/C so far. There were a total of 34 IPS starts in the mission from Nov. 24 1998 to June 30 1999. All other S/C systems operated normally during IPS thrusting. 6) 7) 8) 9) 10)
Table 1: NSTAR/IPS performance characteristics
Figure 7: Diagram of the IPS with the plasma screen removed (image credit: NASA)
The NSTAR/IPS system components like the ion thrusters and the PPU (Power Processing Unit) were manufactured for NASA/GRC by the Hughes Electron Dynamics Division, Torrance, CA; the DCIU (Digital Control and Interface Unit) was built by Spectrum Astro, Inc. The total mass of NSTAR/IPS was about 30 kg; in addition, 82 kg of xenon was stored onboard.
AutoNav (Autonomous navigation):
Objective: evaluation of autonomous on-board orbit determination (determination of S/C location in the solar system and in its flight path) and control capability with the intent to free ground-segment resources. The AutoNav system (a software package including a baseline trajectory, a star catalog, the ephemerides of the asteroids, operations procedures, sequences, etc.) photographs reference asteroids against the background of fixed stars. With the knowledge of time, AutoNav computes the position of the asteroids. By measuring where the asteroids appear relative to the stars, it computes where the S/C must be. It then can project its path to its destination and use its propulsion system to make any course changes that are required. 11)
AutoNav began functioning immediately upon activation of the S/C. The ACS (Attitude Control Subsystem) used a commercial star tracker, AST (Autonomous Star Tracker), to determine its attitude. Then the real-time part of AutoNav correctly provided ACS with the position of the sun as reference so that ACS could turn the S/C to the attitude required to illuminate the solar arrays. About once per week throughout the mission, AutoNav was invoked by the operating sequence to acquire optical navigation images. The system then issues commands to ACS and the integrated camera and imaging spectrometer (MICAS) to acquire visible-channel images, each with one beacon asteroid and known background stars. A heliocentric orbit is computed with a sequence of these position determinations combined with estimated solar pressure, and on-board knowledge of the thrust history of IPS. The trajectory then is propagated to the next encounter target, course changes are generated by the maneuver design element. - Typical AutoNav heliocentric orbit determinations differed from radiometric solutions (references to test AutoNav) by <1000 km in position and <0.4 m/s in the velocity vector. Later refinements improved these values substantially. 12)
Figure 8: The AutoNav software system and interacting system software (image credit: NASA/JPL)
Figure 9: Conceptual view of conventional and autonomous navigation schemes (image credit: NASA/JPL)
MICAS (Miniature Integrated Camera Spectrometer):
MICAS (heritage of Pluto Integrated Camera Spectrometer) was designed and built by a team from USGS, SSG Inc., the University of Arizona, Boston University, Rockwell Science Center, and JPL. The MICAS package, combines the functionality of a framing camera (staring mode) with that of an imaging spectrometer. It features two visible-range imaging channels (APS and CCD - labeled after the detector type used), an ultraviolet imaging spectrometer, and an infrared imaging spectrometer, plus thermal and electronic control. 13) 14) 15)
The CCD array of size 1024 x 1024 pixels is the prime detector, operated in staring mode, for obtaining frame images. The CMOS active pixel sensor (APS) includes the timing and control electronics on-chip with the detector. The UV spectrometer has a spectral range of 80 to 185 nm with 2.1 nm spectral resolution. The infrared spectrometer covers the range from 1200 to 2400 nm with spectral resolution of 12 nm. The imaging spectrometers operate in pushbroom mode.
All four detector systems share a common optical system, namely a telescope with an aperture of 10 cm diameter. With a structure of a highly stable SiC, no moving parts are required. Spacecraft pointing directs individual detectors at the desired targets. The optics system of MICAS was developed by L-3 SSG Tinsley, formerly known as SSG Precision Optronics Inc.
Figure 10: Optical design of MICAS (image credit: NASA/JPL)
Table 2: Some parameters of the MICAS instrument
MICAS serves three functions on Deep Space 1. First, as with all the advanced technologies, tests of its performance to establish its applicability to future space science missions (demonstration of lightweight imaging technology). Second, the visible CCD channel is used to gather images for AutoNav's use. Third, it collects scientific data during this mission at the asteroid flyby and possibly the two comets. The MICAS instrument mass is 12 kg.
Figure 11: Artist's rendering of the MICAS instrument (image credit: NASA)
Figure 12: Illustration of the MICAS instrument (image credit: NASA/JPL)
Except for the UV sensor that became defective during the flight calibration, MICAS was fully operational throughout the entire mission. In particular, the MICAS-CCD channel was extensively used by AutoNav to obtain optical data to update its trajectory knowledge throughout most of the primary mission and the extended mission.
PEPE (Plasma Experiment for Planetary Exploration):
The objective was to assess the performance of this highly integrated, low-mass and low-power instrument. The instrument provided measurements of the three-dimensional plasma distribution over its field of view (2.8 π steradians). PEPE, built by Southwest Research Institute (SwRI) and Los Alamos National Laboratory (LANL), combines multiple instruments into one compact 5.6 kg package. The instrument includes a very low-power-consumption, low-mass microcalorimeter, provided by Stanford University, to help understand the plasma/surface interactions. 16) 17) 18) 19)
During the cruise phase, PEPE measures the solar wind energy spectrum of electrons and ions from 8 eV to 33 keV per unit charge with at least 5% resolution. Instead of using moving parts, it electrostatically sweeps its field of view, achieving a resolution of 45º in azimuth and 5º in elevation. PEPE also measures ion mass in the range of 1 to 500 amu per unit charge at a mass resolution of 5%.
Figure 13: Schematic illustration of PEPE (image credit: NASA/JPL)
PEPE combines two spectrometers into one package. The first spectrometer analyzes the energy and mass of ions and the direction in which they are traveling; the second spectrometer analyzes the energy and angular distribution of electrons coming from all directions in space. Together, the measurements give scientists a better understanding of the state and composition of plasma the spacecraft encounters.
PEPE played three functional roles on DS1. It validated the design for a suite of plasma physics instruments in one small package; it assisted in determining the effects of the IPS on the local plasma environment, including interactions with the solar wind and photoelectrons; and it made scientifically interesting measurements during the cruise and the encounters. This device measured charged particles in space, both electrons and charged atoms, or ions, and made detailed measurements of solar wind.
PEPE made measurements of the solar wind with the IPS on and off. The main result is that SEP (Solar Electric Propulsion) can be used on future missions without interfering with the science payload.
Figure 14: View of the PEPE instrument (image credit: NASA/JPL)
SDST (Small Deep Space Transponder) & KAPA (Ka-band solid-state Power Amplifier):
SDST was built by Motorola; KAPA was developed by Lockheed Martin. The SDST package combines the devices of: receiver, command detector, telemetry modulator, exciters, beacon tone generator, as well as the control functions into one unit with a mass of 3 kg (less than half the mass than would be required without this new technology).
SDST allows X-band uplink and downlink as well as a Ka-band downlink (32 GHz). The Ka-band signal is amplified by a solid-state power amplifier to 2.2 W and a gain of 36 dB with an overall efficiency of 13%. The Ka-band offers a potential link-performance advantage for deep-space communications. With future improvement of ground facilities and spacecraft hardware, assuming similar power efficiencies and spacecraft antenna sizes, Ka-band holds a potential four-fold increase in data rate compared to X-band. 20) 21) 22)
In addition to characterizing the operation of the Ka-band solid-state power amplifier, DS1 provided Ka-band signals for DSN use in verifying systems for acquiring, demodulating, decoding, and processing telemetry as well as in producing 2-way Doppler and ranging data. As the Earth-S/C range increased , certain tests were repeated to assure that the transition through threshold in a selected Ka-band region was observed.
SDST performance monitoring and spacecraft data interfaces:
• Receives commands from the IEM (Integrated Electronics Module)
• Collects analog-engineering status within the system and provides status and performance parameters to the IEM
• The SDST design accommodates interfaces with spacecraft avionics via either a MIL-STD-1553, MIL-STD-1773, or RS422 serial bus, using the 1553 protocol. This design allows future SDST users the maximum flexibility of selecting the system architecture. The DS1 SDST C&DH (Command and Data Handling) communication is via the 1553, and the data interface uses the RS422.
Figure 15: Block diagram of the DS1 RF communication subsystem (image credit: NASA/JPL)
SDST has also the ability to generate beacon signals in Beacon Monitor Operations (Beacon Monitor includes a software that diagnoses the spacecraft's operational condition). The system then transmits one of four tones to indicate to the operations team the urgency of the spacecraft's need for DSN (Deep Space Network) coverage. The four tones correspond to 1) the S/C is healthy and doesn't need any assistance, 2) reporting the occurrence of an unusual but not threatening event permitting network scheduling procedures, 3) alerting the ground that intervention is needed to prevent the loss of important data or to assist in resolving problem, and 4) requiring immediate assistance because the S/C has encountered a mission-threatening emergency. In each case, when tracking is initiated, the data summarization system provides a synopsis of the pertinent S/C data. 23) 24)
Figure 16: Illustration of the SDST unit (image credit: NASA/JPL)
ARAX (Autonomous Remote Agent Experiment):
This is a software package consisting of the following three modules: PS (Planner/Scheduler), EXEC (Smart Executive), and MIR (Mode Identification and Recovery). The objective of this ambitious experiment involves turning over the responsibility (i.e. control) for the S/C to an autonomous agent. The concept involves a new architectural approach, taken by a team from JPL, AMES and the Carnegie Mellon University (Pittsburgh, PA), which uses an agent of the ground team onboard the spacecraft. This remote agent is tested in a restricted case on DS1, in preparation for more ambitious experiments on subsequent flights. The Remote Agent includes an onboard mission manager that carries the mission plan, expressed as high-level goals. A planning and scheduling engine uses the goals, comprehensive knowledge of the spacecraft's state, and constraints on spacecraft operations, to generate a set of time-based or event-based activities, known as tokens, that are delivered to the executive. The executive expands the tokens to a sequence of commands that are issued directly to the appropriate destination on the spacecraft. The executive monitors the response to these commands (through the mode identification and reconfiguration module) and reissues or modifies them if the response is not what was anticipated. Several faults are being simulated during the remote agent experiment on DS1.
On May 17, 1999 the primary S/C command was given over to Remote Agent for three days of S/C operations. In this period, Remote Agent successfully planned DS1 activities on-board and then carried out the plan without ground intervention. The software detected, diagnosed and fixed simulated problems, showing that it can make decisions to keep the mission on track. This Remote Agent capability, a precursor to self-aware and self-controlled robots, will reduce the cost of future S/C operations as computers become “decision-making partners” along with humans.
Figure 17: Conceptual configuration of ARAX (image credit: NASA/JPL)
Note: It turned out that remote agent did have a problem (which, of course, is the reason for testing it!) that prevented it from continuing for the entire three days. The experiment was successful in that the bug was found. However, the bug did not present a risk to the spacecraft, so another experiment was designed and allowed the remote agent to complete all of its test objectives.
LPE (Low Power Electronics):
The LPE experiment was developed at MIT/LL and funded by DARPA. The objective was to validate the performance of electronic devices throughout the life of the mission, with particular interest in the effects of radiation (correlation of the changes with total-dose-radiation measurements). The LPE approach was to observe the properties of test devices where no attempt had been made in either processing or packaging to optimize performance for the radiation environment. The particular interest was to characterize the sub 0.25 µm FDSOI (Fully Depleted Silicon-On-Insulator) baseline process developed at MIT Lincoln Laboratory and verifying that the inherent radiation-hardened qualities of the technology that have been examined through ground testing hold true in the space environment. 25)
The FDSOI parametric testing on DS1 was performed by a board designed to emulate the tasks of a semiconductor parameter analyzer. The sub 0.25 µ FDSOI test chip was mounted on this board. All board components, with the exception of the test chip, were radiation-hardened so that all changes in behavior could be isolated to the test chip.
In addition to monitoring key transistor properties, the LPE also addressed the issue of performance monitoring. By sampling the output frequency of four 97-stage ring oscillators, an evaluation could be made of how the stage delay was affected by the space environment.
The results of the LPE tests have shown that transistor characteristics and performance are minimally affected by the space environment. This insight into the fundamental building blocks of circuit design will prove to be invaluable when creating more complex SOI test circuits for further space qualification.
Figure 18: Photo of the LPE experiment 6µm VME-style Test Board (image credit: NASA/JPL)
PASM (Power Actuation and Switching Module):
PASM was a joint development of JPL, Lockheed and Boeing. The objective is to determine the working performance of PASM (internal test load of up to 40 V and 3 A). The goal is to advance the art of power electronics packaging- namely Lockheed Martin's proprietary HDI (High-Density Interconnect) technology, and Boeing Company's expertise in ASIC (Application Specific Integrated Circuit) design and layout. PASM was cost-shared by both the government and industry. The industry assumed the cost of developing the product, and the government paid for its fabrication and test. 26)
PASM contains two sets of four power switches (quad-switch device). Each of its four stand-alone switches provides the capability to switch power, to isolate faults, and to limit in-rush and fault currents, and supplies voltage and current telemetry. Additionally, it offers the capability for trip time control, di/dt and dv/dt control, and remote on/off control. Each switch can switch anywhere from 3 to 40 V at 3 A maximum and, as a result, can be used in switching the primary as well as the secondary side (conditioned) power. The use of HDI technology for packaging and ASICs for switch control electronics gives PASM a 4 to 1 weight, volume, and footprint advantage over existing hybrid products. It is the advanced packaging technology and utilization of ASICs that makes the PASM unique.
The validation program included flying two PASM modules as a category 3 experiment on DS1. The test program included switching 5 V power to a 1 A resistive load through each of the eight switches (four per module). The switches were also operated in parallel (two at a time) to switch 5 V power to the same 1 A resistive load. Certain electrical design flaws in the switch control ASICs prevented them from operating completely. As a result, the in-rush and fault isolation features of the PASM switches were not tested.
The PASM switches were successfully exercised several times during the mission and showed no performance degradation or inability to function.
Figure 19: Illustration of a single PASM unit (image credit: NASA/JPL)
Figure 20: View of the PASM flight configuration on DS1 (image credit: NASA/JPL)
MFS (Multifunctional Structure):
The MFS experiment was provided by AFRL (Air Force Research Laboratory) and Lockheed Martin Astronautics. Objective: performance test of a multifunctional structure (attached to the S/C bus) consisting of electronic connection systems for embedded devices and the thermal control of a test panel. This new packaging technology combines load-bearing elements with electronic housings and thermal control, thus greatly reducing the mass of S/C cabling and traditional chassis. 27)
Figure 21: Illustration of the MFS concept (image credit: Lockheed Martin)
Note: The first MFS technology demonstration was flown on the STEX (Space Technology Experiment) mission of NRO (launch Oct. 3, 1998). Other missions followed like: DS1 (Deep Space 1, launch Oct. 24, 1998) of NASA, EO-1 (Earth Observing, launch Nov. 21, 2000) of NASA, and the STRV-1d (Space Technology Research Vehicle-1d, launch Nov. 16, 2000) mission of DERA, UK. At the start of the 21st century, MFS is a proven and accepted technology used in virtually all space missions. 28)
SCARLET-II (Solar Concentrator Array with Refractive Linear Element Technology):
SCARLET-II was built by AEC-Able Engineering Co., of Goleta, CA, with BMDO (Ballistic Missile Defense Organization) funding. SCARLET is the high-power solar array of DS1; it uses cylindrical silicone Fresnel lenses to concentrate sunlight onto a strip of photovoltaic cells and acts to protect the cells. Each array is composed of four panels of size 160 cm x 113 cm. The multijunction GaInP2/GaAs/Ge photovoltaic cell modules are interconnected in series to produce about 2500 W (at nominally 90 V) at the beginning of the mission. - The dual-junction cells achieved an average efficiency of 22.5% established during in-flight analysis. 29) 30) 31) 32) 33)
The key measure of performance, specific power, demonstrated on DS1 is at the state-of-the-art (45 W/kg) and can easily be increased with now proven design enhancements and/or size increase.
Figure 22: Photo of the DS1 Scarlet (Wing 1 of 2) on deploy rail (image credit: AEC-Able Engineering Co. Inc.)
Legend to Figure 22: Wing size: 5.20 m x 1.42 m; panel dimensions: 1.14 m x 1.60 m (4 panels per wing); array power = 2.5 kW (1 AMO); wing mass = 27.7 kg.
Major events and status of extended mission
• On July 29, 1999, DS1 successfully performed a close flyby of asteroid 9969 Braille using the AutoNav system. At about 27 km separation, it was by far the closest flyby of an asteroid ever attempted. - Note: The asteroid, discovered in 1992, was only recently (1999) named in honor of Louis Braille (1809-1852), the Frenchman, who invented the alphabet for the blind. 34) 35)
• Two months after the end of its extremely successful primary mission, the Star Tracker of DS1, responsible for the spacecraft's orientation, ceased operating. NASA decided it could afford to go ahead with a risky extended mission for a comet encounter. Thereafter, JPL engineers devised a way (a long-distance rescue mission) to restore the spacecraft's sense of direction by writing new computer programs to use the camera (MICAS) as a substitute Star Tracker. The rescue also involved developing new operational procedures (new methods of flying the spacecraft) that went along with the new software. The challenging task was completed in June 2000 to resume thrusting in time to give DS1 a chance to encounter the comet Borrelly in September 2001.
• On Sept. 22, 2001, DS1 flew by the comet at a speed of 16.5 km/s (DS1 came within 2200 km of 19P/Borrelly's 10 km long core, transmitting 30 black-and-white images of the coma. DS1 instruments also catalogued the gases issuing from the comet's coma, which included water vapor and carbon monoxide, and measured their interaction with the solar wind as well as their infrared radiation characteristics. One surprising observation was that the gases were projected asymmetrically from the coma.). 36)
• There were formidable obstacles prior to achieving this feat. For one, the camera's field of view is 100 times smaller than the Star Tracker's. Also, while the Star Tracker could estimate the probe's orientation in space four times every second, MICAS produces a computer file that takes more than 20 seconds to transfer to the computer for analysis. Updates from the camera would be at least 80 times slower than the original guidance system.
• After all preparations were completed and tested successfully, the ion engine was turned on June 21, 2000 (after a hiatus of 7 months) for tests of operating it with the new control system. On June 28 thrusting began in the direction required to reach comet Borrelly. - At launch, the plan for a possible extended mission was to go to Borrelly. After the spacecraft had been flying for about half a year, JPL decided that the mission was going so well and the advanced technologies were working so well that comet Wilson-Harrington could be added to the extended mission proposal. When the star tracker failed, project management reverted to the earlier plan.
Figure 23: Artist's view of the DS1 encounter with comet Borrelly (image credit: NASA/JPL)
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The information compiled and edited in this article was provided by Herbert J. Kramer from his documentation of: ”Observation of the Earth and Its Environment: Survey of Missions and Sensors” (Springer Verlag) as well as many other sources after the publication of the 4th edition in 2002. - Comments and corrections to this article are always welcome for further updates.