Date: Wed Feb 23 1994 12:08:00 Subj: Clementine The DSPSE mission involves approximately t

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Date: Wed Feb 23 1994 12:08:00 From: Royston Paynter Subj: Clementine The DSPSE mission involves approximately two months in lunar orbit and then departure on a trajectory that will cause the spacecraft to pass by the asteroid 1620 Geographos at close range. The time in lunar orbit provides an opportunity to collect lunar images and data for scientific investigation. The close range flyby of 1620 Geographos provides an opportunity to collect images and data from the asteroid for scientific investigation. The mission is composed of the following operational phases: o Low Earth Orbit (LEO) o Lunar Transfer Trajectory: Translunar Transfer Injection (TTI) to Lunar Insertion o Lunar Orbit o Geographos Transfer Trajectory: - Lunar Departure to Lunar Swingby - Lunar Swingby to Geographos Pre-Flyby o Geographos Pre-Flyby and Flyby o Post-Flyby Operations o Post-Mission Support Lunar Orbit Operations: The lunar mapping phase of the DSPSE mission is planned to last 70 days and will include 57 days of mapping of the lunar surface with the imaging sensors. The first five days in lunar orbit will be used to establish the lunar mapping orbit and for pre-mapping mission readiness activities including calibration of the sensors, autonomous position estimation tests, and imaging special areas of interest on the lunar surface. The next 27 days will be devoted to lunar mapping. After the first lunar sidereal day of mapping, approximately one earth day will be used to change the latitude over which periselene occurs and the next 27 days will be devoted to the last lunar sidereal day of mapping. During the lunar mapping phase, mapping will be temporarily interrupted for orbit adjustments and may be interrupted to divert the sensors from lunar mapping to observe higher priority targets of opportunity. The final 8 days following mapping operations will be used for preparations for injection into the Geographos transfer trajectory, autonomous position estimation tests, and possible observations of targets of opportunity. The targets of opportunity may involve a change of orbit if adequate fuel reserves remain. Another possibility, without changing the orbit, is to divert the sensors from nadir pointing to observe targets of opportunity. The Lunar Orbit On 21 February 1994, the lunar injection maneuver will place the spacecraft into a preliminary lunar orbit passing over the lunar south pole. The orbit injection will be accomplished with two burns using an ~ 80%/20% split in total Fv with time in between for orbit determination. After the first burn the spacecraft will be in an orbit with an approximate period of 8 hours. The inclination of the orbit will be 90!11! with reference to the lunar equatorial plane. DSN range and range-rate and Pomonkey range-rate and spacecraft accelerometer data will be used by GSFC to calculate the final orbit shaping maneuver parameters. After approximately 24 hours (2 orbits) the second burn will place the spacecraft in the initial mapping orbit with a periselene altitude of 425125 km occurring at -27! to -30! lunar latitude (TBD) on the sunlit side of the moon with a 5 hour orbital period to allow for 131.5 orbits per sidereal month. This geometry allows mapping the southern polar region at a lower altitude. Orbit shaping will be performed, as required, to adjust periselene and orbital period to meet mission requirements. On 25 March 1994, after approximately 32 days, mapping operations will be temporarily suspended for approximately one day to allow the mapping orbit to be changed so periselene will occur over +30! to +27! lunar latitude. This maneuver will allow mapping of the northern polar region at a lower altitude for the second month of mapping. On 29 April 1994, two orbit adjust burns will be made to change the periselene so it occurs over +48! latitude. This orbit will be the departure orbit. On 3 May 1994, a lunar orbit departure RCS burn will be performed to leave the Lunar orbit and to begin the trans-Geographos trajectory. Lighting Conditions The lunar mapping effort will consist of topographic imaging, altimetry and multispectral imaging for mineral identification. The best data for the lunar mineral mapping mission presently considered will be obtained if the solar phase angle is less than 30!. The solar phase angle is defined as the angle between the vector to the Sun and the vector to the spacecraft from a point on the Moon's surface. To maximize the time period in which the solar phase angle is within 30!, the plane of the lunar orbit will contain the Moon-Sun line half way through the two-month lunar mapping period. Therefore, insertion into the lunar orbit has been selected so that, as the Moon-Sun line changes with Earth's motion about the Sun, the Moon-Sun line will initially close on the orbital plane, and then lie in the orbital plane half-way through the mapping mission. Assuming a two-month mapping mission and several days checkout time in lunar orbit prior to beginning mapping, the angle between the Moon-Sun line and the orbital plane would close for approximately five weeks before becoming zero. The current lunar planning trajectory has this angle set at 28.3! at insertion into lunar orbit, so that at the beginning of mapping (5 days after insertion) it will be 23.6!. Half way through mapping it will be -2.4! and at the completion of mapping, it will be -30.3!. At lunar departure (70 days after insertion) this angle will be -40.3!. Spacecraft Activities and Tests A more detailed description of specific experiment activities is contained in section 6.4.5. The detailed experiment operations are described in the Appendices. The main spacecraft activities and tests to be performed during the 70 day lunar mission include: Pre-Lunar Mapping Checkout Activities: The initial 5 days will be devoted to establishing the proper mapping orbit, autonomous position estimation tests, observing special areas of interest on the lunar surface, and preparations for lunar mapping. These preparations include: - Verify the gain settings and exposure times (integration time settings) for each camera and assessing the ability of the on-board software to control these settings; - Determine the range of altitudes for which the laser ranging system can measure the range to the lunar surface; - Assess jitter effects from spacecraft pointing activities, filter wheel movement, solar array movement, and cryocoolers in a lunar mapping scenario; - Assess orbit determination accuracy and error analysis; - Perform Autonomous Position Estimation Experiment activities. This will involve imaging the lunar limb from the dark side of the moon. - Perform Autonomous Operations and Scheduling including testing of the auto-sequence code; - Assess on board maintained pointing vectors to the center of the Moon, Sun, Earth, and each of the ground stations; - Rehearse lunar mapping operational readiness : -- Command script generation, verification, uplink, and execution; -- Collection of imaging data; -- Real-time downlink of image data during data collection; -- Downlink of image data stored on the Solid State Data Recorder (SSDR); -- Data transfer, processing, and display. - Collect high quality images and altimetry and downlink to the ground. Lunar Mapping Activities: The formal collection of lunar mapping data is planned to take approximately 55 days or two lunar sidereal days (1 lunar sidereal day is approximately 27.3 earth days)]. The UV/Visible (UV/Vis) camera with 5 filters and the near infrared (NIR) camera with 6 filters will each collect overlapping images. Each filter is used along the satellite's ground track such that complete mapping of the lunar surface for each of the filters of each camera will be accomplished during the mapping operation. The High Resolution (HiRes) camera will make a continuous string of overlapping images with TBD filters along the track of the satellite on the sunlit side of the Moon. The laser ranging system will be used when the spacecraft altitude is below 500 km (TBD) at a pulse rate of 1 Hz (TBD) to preclude exceeding its thermal constraints. The LWIR camera will be used to make images of areas of interest along the spacecraft's path, especially near the terminator on the sunlit and dark sides to collect temperature change information. Since this requires terminator motion along the lunar surface, off-nadir pointing will be required. Complete mapping of the lunar surface will not be accomplished with the HiRes camera, the laser ranging system or the LWIR camera. Autonomous position estimation and orbit determination tests, and autonomous operations scheduling tests flowing from the position estimation and orbit determination tests, will be conducted during all portions of the lunar orbit phase. A goal is to verify the accuracy of the autonomous position estimation and orbit determination and operations scheduling functions during the earlier portion of the lunar orbit phase and then let the spacecraft schedule and execute operations for one or more lunar orbits during the later portion of the lunar orbit phase. Preliminary Operational Scenario: The formal collection of lunar mapping data is planned to take approximately 55 days or two lunar sidereal day (1 lunar sidereal day is approximately 27.3 earth days)]. The UV/Visible (UV/Vis) camera with 5 filters and the near infrared (NIR) camera with 6 filters will each collect overlapping images. Each filter is used along the satellite's ground track such that complete mapping of the lunar surface for each of the filters of each camera will be accomplished during the mapping operation. The UV/Vis and NIR cameras need to be turned on 10 minutes before imaging. The NIR cryocooler needs to be on 30 minutes before imaging. In successive orbits, the UV/Vis and NIR camera images will be taken alternatively from -90! to +60! and -60! to +90! latitude, when periselene is at -30! latitude, and -60! to +90! and -90! to +60! latitude, when periselene is at +30! latitude. This mapping strategy is designed to save power and storage space since there is no need to take images of polar regions on consecutive orbits. The integration times and gains will be varied as a function of latitude and may be computed on board. The time to change filters and to dampen is assumed to be 200 -250 ms. The UV/Vis camera has a stray light exclusion angle of 40! (full angle). The UV/Vis camera, along with the star trackers will also be used to collect data for autonomous orbit determination tests by obtaining lunar limb and star images. There is also a requirement to obtain frequent sensor calibration data for all the mission sensors. These tests will be taken for TBD minutes each orbit as the timeline allows. The High Resolution (HiRes) camera will make a continuous strip of overlapping images with up to 4 filters along the track of the satellite on the sunlit side of the moon. Images using the HiRes camera will be taken -90! to +90! latitude. The camera integration time and gains will be varied with the latitude and may be computed on board. The time to change filters and to dampen is assumed to be 200 -250 ms. The laser ranging instrument (LIDAR) will be used whenever the altitude is less than TBD km. The maximum usable altitude will be determined while in Lunar orbit, but will not exceed 640 km. During initial lunar orbit operations, tests will be conducted to determine the maximum altitude from which effective altimetry can be performed. The laser ranger will be turned on TBD minutes prior to use. The LIDAR electronics will be turned on 10 minutes before use and the heaters will be turned on 15 minutes prior to use to bring laser transmitter temperature up to +25 degrees. The laser pulse rate will be limited to 1 Hz to keep temperatures in range. The Long Wavelength Infrared (LWIR) camera will be used to take images of thermal gradients occurring at the terminators. The cryocooler for the LWIR must be on 30 minutes prior to imaging. The LWIR camera electronics will be on 10 minutes prior to imaging. Images will be taken 110! of each terminator over the poles and the camera will be turned off between usages, but the cryocooler will remain on between North and South pole image sequences. As with the NIR, UV/Vis and HiRes cameras the integration time and gain will be varied with the latitude and may be computed on board. During lunar mapping, the star trackers will not be used to collect scientific data, but will be used to establish the attitude of the spacecraft. To meet the high accuracy pointing requirements, the spacecraft attitude will be updated every 10 seconds during mapping. Star tracker images will be processed on the R3000 image processing computer and used to update the attitude of the spacecraft. The star trackers have a full angle solar exclusion angle of 63! x 75! in order to resolve stars. One star tracker will have solar exposure during each lunar orbit and will be powered off for this time. Solar panel auto tracking will not be inhibited during imaging. The exact solar panel position management procedures to minimize jitter is TBD and will be determined during simulations and verified during the first 5 days of lunar orbit. During lunar imaging, the spacecraft will use nadir pointing to meet sensor requirements for imaging. The best method will be finalized during the initial 5 day operational readiness phase. Attitude changes and attitude updates must be scheduled with the camera image sequences and positioning of the solar panels. Attitude commands will be made to the guidance software, which will generate attitude control parameters for the desired attitude and supply these parameters to the ACS software. Attitude parameters will be generated based on any attitude constraints, primary pointing requirements (nadir or earth pointing), and secondary pointing requirements (maximizing solar incidence angle on solar arrays). The ACS software will compare the guidance generated attitude parameters with the current attitude and generate the required commands to slew to the desired attitude. The reaction wheels or thrusters will be used for the slew depending on the amount of slew and time available for the slew. Slews using the reaction wheels will take up to 10 (TBD) minutes maximum. After lunar mapping sequences and autonomous position estimation and sensor calibration tests are complete for an orbit, the spacecraft will dump data to the ground using the high gain antenna. This dump will take between ~120 to 130 minutes to completely downlink the data from the SSDR. Engineering data will be stored and dumped along with the image data using the high gain antenna. Engineering and limited image downlink via the omni antennas during lunar mapping will be possible when orbital geometry and ground station visibility permit. Post Mapping Activities: During the 8 days following the completion of the lunar mapping there will be time for special observations of targets of opportunity. These observations will provide images of lunar surface features and man-made objects on the Moon at the Apollo, Surveyor, and Russian landing sites. There will be targets that will be imaged without changing the orbit. Other targets of opportunity may involve a change of orbit for observations which will be done if adequate fuel remains for the Geographos transfer and flyby. See section 6.4.7 for further details. Prior to lunar departure there will be two orbit adjust maneuvers to prepare for the lunar deorbit maneuver. This will change the periselene so that it occurs at +48! latitude. Pomonkey will supply range-rate and the DSN sites will supply range and range-rate data for orbit determination to GSFC who will compute the state vector and supply it to the DMOC. The TAMP will verify parameters, develop command sequences, and verify these sequences using the OTB spacecraft simulator. The Lunar Orbit Departure will be fully rehearsed to acquaint operations personnel with the activities and timing of the departure RCS burn and to verify the activities and sequences can be smoothly executed. Ultraviolet/Visible (UV/Vis) Camera The UV/Vis sensor is a CCD video camera with an 8-bit digitization of the array data. The actual ground resolution from 400 km (TBD) altitude is between 79 to 106 meters depending on the actual amount of jitter. Six bandpasses are defined by filters in a six position filter wheel. One of the six filter wheel positions will be required for a very wide bandpass filter, 400 to 950 nanometers (nm), so virtually the entire sensor bandpass can be used to allow early identification of Geographos and closed-loop tracking control. The five remaining filter bandpasses were specified by the NASA Science Advisory Committee (SAC) and SDIO. The UV/Vis camera will be mounted so the 4.2! dimension of its field of view is in the spacecraft's velocity direction (X-axis - along track) for the lunar mapping phase and the 5.6! dimension of its field of view is perpendicular to the spacecraft's velocity (Y-axis - cross track). For temperature stabilization, the UV/Vis camera electronics must be turned on for 10 minutes before it can be used to collect image data. Radiation exposure could damage the UV/Vis; however, the effects can be corrected on the ground. Near Infrared (NIR) Camera The NIR sensor is a cooled video camera with an 8-bit digitization of the array data. The actual ground resolution from 400 km (TBD) altitude is between 112 to 128 meters depending on the actual amount of jitter. The NIR has a six position filter wheel with the filters as shown in Table 6.2-2 selected by the DSPSE SAC and SDIO. The NIR sensor has a mechanical cooler (cryocooler) to bring the CCD array to a sufficiently low temperature for operation. For temperature stabilization, the NIR camera's cryocooler must be turned on for 30 minutes and the camera electronics must be turned on for 10 minutes before it can be used to collect image data. Long Wavelength Infrared (LWIR) Camera The LWIR sensor is a cooled video camera with an 8-bit digitization of the array data. The actual ground resolution from 400 km (TBD) altitude is approximately 43 meters and is fairly independent of the amount of jitter. The target temperature range for the LWIR is 250 to 400 K. The sensor does not have a filter wheel. The LWIR sensor has a mechanical cooler (cryocooler) to bring the CCD array to a sufficiently low temperature for operation. For temperature stabilization, the LWIR camera's cryocooler must be turned on for 30 minutes and the camera electronics must be turned on for 10 minutes before it can be used to collect image data. High Resolution (HiRes) Camera The HiRes camera is the imaging portion of the LIDAR system. The HiRes sensor uses a frame transfer CCD with a micro-channel image intensifier and uses an 8-bit digitization of the array data. With the current jitter constraints, the HiRes camera is expected to give a maximum ground resolution at lunar periselene of 13 m to 30 m (short to long integration time.) Six bandpasses are defined by filters in a six position filter wheel. One of the six filter wheel positions will be required for a very wide bandpass filter, 400 to 750 nm, so virtually the entire sensor bandpass can be used to allow early identification of Geographos and closed-loop tracking control. Another of the six filter wheel positions will be opaque to protect the intensified camera from high light input when it is not in use. The four remaining filter bandpasses were specified by the DSPSE SAC and SDIO. Radiation exposure could damage the HiRes camera, however, the effects can be corrected on the ground. Solar illumination could also damage the HiRes camera, so imaging should not take place if the Sun is within 2.38! from its optical axis, and the opaque filter should be selected if the sensor is not being used for active imaging. For temperature stabilization, the high resolution camera electronics must be turned on for 10 minutes before it can be used to collect image data. Laser Ranging System (LIDAR) The laser ranging system comprises a Nd:YAG laser and a ranging receiver. The laser emits 180 millijoules per pulse at 1064 nanometers with a beam divergence of 500 microradians. The ranging receiver is an avalanche photo diode with a 1 milliradian full-angle field of view. The laser ranging system is able to resolve range measurements to 40 meters. Ranging activity is desirable while the kick motor recedes from the spacecraft during the lunar transfer trajectory, during the flyby of the asteroid, and during any low altitude portions of the lunar orbit phase. The laser ranging system design has been modified to permit range measurements to ranges of 500 km above the lunar surface. The laser has a maximum pulse repetition rate of 8 pulses per second, but is not expected to sustain a pulse repetition rate greater than 1 pulse per second without thermal problems. For temperature stabilization, the laser heater must be turned on for 15 minutes before it can be used for ranging. Star Trackers The star trackers are star imaging sensors used for 3-axis Rlost in spaceS attitude determination. The star trackers consist of a S-20 photocathode and a full frame CCD video camera with an 8-bit digitization of the array data. Associated with the star tracker is a set of algorithms and a catalog of selected stars which can be used with data from the star tracker to determine its attitude to an accuracy of 150 microradians in pitch and yaw and 450 microradians in roll. With a 150 millisecond integration time, the star trackers can detect stars down to magnitude 4.5. For reliable operation, sources such as the Sun, reflected sunlight, or bright limbs must be excluded from a 63! by 75! region about the star trackerUs optical axis. Staring at the Sun may damage the star trackers, therefore, the star tracker optics must not be exposed to direct sunlight (sun within the 28! by 42! FOV) for a period exceeding 3 (TBD) minutes, operating or non-operating. The two star trackers will each have a different line of sight selected to allow at least one of the star trackers to determine the spacecraft's orientation while lunar mapping is being conducted and during all other phases of the mission. One star tracker will have solar exposure during each lunar orbit and will be covered and powered off during this time. There is also a potential for radiation exposure damage to the star trackers which may require onboard processing to correct. For temperature stabilization, the star tracker electronics should be turned on 10 minutes before images are obtained for downlink. Star tracker imaging for attitude determination does not have to wait 10 minutes and can be used right away. Science Objectives: The science objective of the mission is to obtain data useful for scientific investigations. These investigations involve using the DSPSE mission sensors to image the lunar surface and the asteroid Geographos. The operational experiments for these events were discussed above. This section discusses the scientific reasons for performing these observations. The DSPSE mission will provide an abundance of information about the surface morphology, topography, and composition of both the Moon and Geographos, providing an insight to their history and processes that have shaped that history. This information will be used to address fundamental questions in lunar science and will contribute to significant advances toward deciphering the complex story of the Moon. The DSPSE mission will also permit a first-order global assessment of the resources of the Moon and provide a strategic base of knowledge upon which future missions to the Moon can build. Lunar Mapping The pressing need for global mapping of the Moon, by a variety of remote-sensing techniques, has been stressed for the last 20 years. The DSPSE mission begins this task allowing a global digital image model (DIM) of the Moon to be developed. DSPSE lunar mapping will provide improved resolution allowing for more detailed geologic mapping and will obtain improved spectral coverage as well as improved spectral resolution over any previous lunar observations. While the Galileo spacecraft provided spectacular multi-spectral images of the lunar surface before leaving the Earth/Moon system, the DSPSE mission offers many significant advances relative to the Galileo data. o The pixel resolution will be at least 10 times better than Galileo's, providing improved resolution for unit mapping. o The HiRes camera images offers up to 100 times better resolution, allowing for detailed geologic mapping. o Using the various sensors, improved spectral coverage (up to 2.8 microns) will be obtained. o With the increased number of filters on the sensors, improved spectral resolution will be possible allowing improved distinction between olivine, pyroxenes, and plagioclase. The altimetry obtained by the laser ranger, will augment the DIM by providing a set of topographic profiles for the mid-latitude band of the Moon. The DSPSE data when tied to the Apollo data, will permit knowledge of the true positions of lunar surface features to within a few hundred meters. Maps of the Moon made from DSPSE data will enable studies of regional history and permit the processes of volcanism, tectonism, and impacts that have shaped lunar history to be deciphered. The DSPSE data will allow several key unresolved scientific issues to be addressed: o Character and evolution of the primitive lunar crust. o Thermal evolution of the moon and lunar volcanism o The impact record and redistribution of crust and mantle materials o Distribution of potential resources. From the combined UV/Vis and NIR camera images, a global color map will be formed that can be interpreted in terms of rock types. At a minimum, it will be possible to recognize and discriminate between the absence of mafic minerals (pure feldspar) and the presence of orthopyroxene, clinopyroxene, and olivine, as has been done for the near side of the Moon from Earth-based data. Thus on a global basis, the distribution of anorthosite, RnoriticS rocks, olivine-bearing rocks (dunites and troctolites), and gabbros will be able to be distinguished. For mare deposits, visible color mapping can classify the mare in terms of titanium abundance, and element that can be used to estimate the distribution of solar wind hydrogen, an important lunar resource. Combined with the knowledge of cratering and the use of basins as probes of the crust, these data will permit the composition and petrologic structure of the crust to be reconstructed in three dimensions. The question of the existence of a magma ocean, the nature of Mg-suite magnetism, the history and extent of ancient KREEP and mare volcanism, the compositional diversity of mare units, and the effects of cratering on the composition of the lunar surface can be addressed. Topographic data from the laser ranger combined with spectral information will allow the dynamics of large impacts ,e.g., the problem of depth of excavation for basin-sized impacts, to be modeled. The high-resolution images from the HiRes camera will allow the surface processes and compositions to be studied in greater detail. Many mare units display significant heterogeneity, and color imaging from the DSPSE HiRes camera images can map different color units, some of which are perhaps related to individual mare flows. Images of crater walls and central peaks can not only provide high-resolution compositional data, but permit a better understanding of the geological setting and processes that have affected given regions, information that may prove critical to the proper interpretation of the regional compositional information. Finally, the high-resolution imaging can be used to make detailed geological studies of the areas of high scientific interest. Finally, the tracking information received as part of the normal orbit determination tasks during the period in lunar orbit will also be used to provide a more accurate model of the moonUs gravitational potential. Asteroid Flyby Objectives: Asteroid flybys are necessary for the characterization of multiple targets to address issues of asteroid diversity. Flybys cannot address fundamental questions of elemental composition (for link to meteorites) nor of internal structure - both of these will require rendezvous missions. The DSPSE mission will set the stage for NASAUs Discovery Program Near-Earth Asteroid Rendezvous (NEAR) Program. The NASA science objectives for the Geographos flyby are exploration and mapping. Exploration is the key characteristic - no Near Earth Asteroid (NEA) has ever been investigated close-up before. The mapping will provide insight into the geological and thermal processes characteristic of this type of asteroid. Other information to be gained by the flyby and resultant images will be: o Volume/shape determination o Spin state determination o Multi-spectral imaging to infer compositional heterogeneity o Investigation of regolith with thermal imaging of the dark side. The DSPSE data will allow several key unresolved scientific issues to be addressed: o Relationship of NEAs to main belt asteroids, comets, meteorites, and the planetesimals that were the building blocks of the terrestrial planets (need diversity). o Characteristics of NEAs which have influenced EarthUs geological/biological evolution o Potential of NEAs for sample return and resource utilization. o Surface processes of small objects with weak gravity fields. ================================================================ Lou Wheatcraft, Phone: (713)280-1892; Fax: (713)283-7903 E-Mail: lsw@bonnie.jsc.nasa.gov ================================================================

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