Clementine Mission

                Information from the CD-ROM VOLINFO.TXT file                                
                    Assembled from many sources by:                           
                               Eric Eliason                                    
                     Planetary Data System - Imaging Node                     
                         Branch of Astrogeology                               
                     United States Geological Survey                          
                            October 1, 1994                                   
                    TABLE OF CONTENTS                                         
1. INTRODUCTION                                                               
2. CLEMENTINE EDR IMAGE ARCHIVE                                               
3. MISSION TIME LINE                                                          
4. THE SCIENTIFIC PAYLOAD                                                     
5. LUNAR ORBITAL DYNAMICS                                                     
6. INITIAL SCIENTIFIC RESULTS                                                 

A separate text file containing the times and locations of the orbits, 
and the CD-ROM volumes containing the imaging data for a given orbit.
1. Introduction                                                               
During the past decade, the Department of Defense and the Ballistic           
Missile Defense Organization, BMDO (formally the Strategic Defense            
Initiative, SDIO) of the US Department of Defense (DOD) have invested         
heavily in space technology, focused on the development of lighter, more      
cost effective components and systems for spacecraft.  In May of 1990,        
the NASA Discovery Science Working Group endorsed a study of                  
incorporation of SDIO developed technology into civilian scientific           
missions. In September 1990, NASA Administrator Richard Truly formally        
inquired of Deputy Secretary of Defense Donald Atwood as to the               
possibility of utilizing the advanced, lightweight technologies being         
developed by the Department of Defense in a joint space exploration           
mission.  After a six month effort by NASA, DOD, and industry, it was         
concluded that a mission to a near-Earth-asteroid was desirable and           
feasible from both organizations' standpoint. In addition to the              
scientific value of an asteroid flyby and of the potential benefit of         
technology transfer between the agencies for NASA, there was clear            
benefits to DOD. In early 1992 the mission was incorporated into the          
Ballistic Missile Defense Organization Sensor Integration Program. The        
mission was subsequently expanded by two months to include mapping of         
the Moon to demonstrate spacecraft and sensor performance prior to the        
demanding asteroid flyby mission. The interstage adapter of the               
spacecraft, housing the solid rocket motor, was also designated as a          
platform to remain in highly eccentric earth orbit as a radiation             
experiment platform and additional sensor target.                             
The Clementine spacecraft was built at the US Naval Research Laboratory       
in Washington, DC, and carried sensors, attitude control systems and          
software designed and built by the Lawrence Livermore National                
Laboratory (LLNL). The USAF supplied advanced lightweight composite           
structures and the launch vehicle, a Titan II B refurbished ICBM.             
Several other organizations were involved, especially NASA with               
communications support, through the Jet Propulsion Laboratory's (JPL)         
Deep Space network, and orbit determination and operations support from       
both the Goddard Space Flight Center and JPL. Supporting these                
laboratories were scores of industrial contractors, both large and            
The spacecraft consists of an octagonal prism about 2 meters high. A          
110-pound thruster for delta-V maneuvers is on one end of the prism and       
a high-gain fixed dish antenna is on the other end. The spacecraft            
propulsion system consists of a nonpropellant hydrazine system for            
attitude control and a bipropellant nitrogen tetraoxide and monomenthyl       
hydrazine system for the maneuvers in space. The bipropellant system had      
a total capability of about 1.9 km/s with about 550 m/s required for          
lunar insertion and 540 m/s for lunar departure. The power system             
consists of a gimbaled, single axis, GaAs/Ge solar array providing a          
total spacecraft power of 360 watts at 30 Vdc, with a specific power of       
240 w/kg, based on lightweight construction. Two arrays of the rotating       
solar panels protrude from opposite sides; by rolling the spacecraft and      
rotating the panels, full solar illumination of the panels would be           
achieved. The solar array was used to charge a 15 A-h, 47-w hr/kg, Nihau      
common pressure vessel battery. The total mass of the spacecraft in the       
launch configuration was 1690 kg, with most of the weight in the solid        
rocket motor required for translunar insertion. The spacecraft dry mass       
is about 227 kg, with a roughly equal mass for liquid fuel. This weight       
was achieved by incorporating many of the previously discussed                
lightweight technologies.                                                     
The main instrumentation on Clementine consists of four cameras, one          
with a laser-ranging system. The cameras include an ultraviolet-visual        
(UV-VIS) camera, a long-wave infrared (LWIR) camera, the laser-ranger         
(LIDAR) high-resolution (HiRes) camera, and a near-infrared (NIR)             
camera. The spacecraft also has two star tracker cameras, used mainly         
for attitude determination, but they also serve as wide-field cameras         
for various scientific and operational purposes. The sensor package has       
a mass of 8 kg and represents the state-of-the-art at the initiation of       
the Clementine program. The sensors are all located on one side, 90           
degrees away from the solar panels. Clementine has 12 small attitude          
control jets that were used to orient the spacecraft to point the             
cameras to desired targets. The attitude control system includes the two      
star tracker cameras and two lightweight inertial measurement units,          
based on a ring laser gyro and an interferometric fiber optic gyro.           
During initial deployment at low-earth-orbit, the full size spacecraft        
was 3-axis stabilized. The spacecraft was 3-axis stabilized in lunar          
orbit via reaction wheels, with a precision of 0.05 Deg. in control and       
0.03 in knowledge.                                                            
The spacecraft data processing was performed by 3 computing systems. A        
MIL-STD-1750A computer with a capacity of 1.7 million instructions per        
second was used for save mode, attitude control system, and housekeeping      
operations. A reduced instruction set computer (RISC) 32-bit processor        
with 18 million instructions per second was used for image processing         
and autonomous operations. The Clementine mission represents the first        
long duration flight of a 32 bit RISC processor. Also incorporated is a       
state-of-the-art image compression system provided by the French Space        
Agency CNES. A data handling unit with its own microcontroller sequenced      
the cameras, operated the image compression system, and directed the          
data flow. During imaging operations, the data were stored in a 3 kg, 2       
Gbit dynamic solid state data recorder and later transferred to the           
ground stations using a 128 kb/s downlink. The spacecraft was commanded       
from the ground using a 1 kb/s uplink from the NASA Deep Space Network        
and DOD stations. Demonstration of autonomous navigation including            
autonomous orbit determination was a major goal of the Clementine             
mission. Autonomous operations were conducted in lunar orbit.                 
2. Clementine EDR Image Archive                                               
The Clementine EDR (Engineering Data Record) Image Archive will contain       
approximately 100 CD-ROM volumes for distribution to the science              
community. Each volume will contain approximately 3 lunar orbits but may      
vary in the number of orbits depending on the size of each orbit. The         
first volume in the series begins with orbit 32--the start of the lunar       
mapping phase of the mission. The Low Earth orbit, and earth phasing          
loop image data will be generated on the volumes at the end of the            
volume series.                                                                
Additional documentation exists on the CD-ROM volume series that              
describe the organization and content of the Clementine EDR Image             
Archive. Please see the EDRSIS and ARCSIS documents located in the            
[document] directory. The EDRSIS document provides a detailed                 
description of the format of the image files that are contained in the        
archive.  The ARCSIS document describes the organization of the               
3. Mission Time Line                                                          
Clementine was launched on January 25, 1994 from Vandenberg Air Force         
Base aboard a Titan IIG rocket. After two Earth flybys, lunar insertion       
was achieved on February 19-th. Lunar mapping took place over                 
approximately 2 months in two systematic mapping passes over the Moon         
(See Lunar Orbital Dynamics section for more details on systematic            
mapping and Appendix I showing a table of the Lunar orbit time lines.)        
Table 1 provides a synopsis of the mission time line.                         
After successfully completing the Lunar mapping phase of the mission,         
Clementine suffered an on-board malfunction at 9:39 AM EST, Saturday,         
May 7, 1994. The result of the malfunction prevented Clementine from          
performing the planned close flyby of the near-Earth asteroid Geographos      
scheduled for August 1994. Preliminary analysis traced the cause of the       
malfunction to the on-board computer which controls most of the               
satellite's  systems including the attitude control thrusters. The            
computer activated several thrusters during a 20 minute telemetry             
interrupt with the ground station, thus depleting all the fuel in the         
Attitude Control System (ACS) tanks. With the  depletion of the ACS           
tanks, the spacecraft was left spinning at 80 revolutions per minute          
with no remaining fuel left to despin the spacecraft.                         
           Table 1 - Mission Time line Summary                                
 3/24/1992   -  Clementine Design Begins                                      
 1/25/1994   -  Spacecraft launched at Vandenberg Air Force Base              
 2/ 3/1994   -  Departure from low Earth orbit following an                   
                eight day checkout period                                     
 2/19/1994   -  Following two phasing loop orbits, Clementine                 
                injected into lunar orbit. Lunar orbit shake down             
                period begins.                                                
 2/26/1994   -  Start of systematic mapping for first cycle.                  
 3/26/1994   -  First mapping cycle complete, start of second cycle           
 4/21/1994   -  Second mapping cycle complete, begin special                  
 5/ 5/1994   -  Leave Lunar Orbit                                             
 5/ 7/1994   -  Clementine satellite suffered on-board malfunction            
                that prevents planned close fly-by of the near-Earth          
                asteroid Geographos                                           
 8/31/1994   -  Planned Geographos Fly-by (not accomplished)                  
4. The Scientific Payload                                                     
UV-Visible Camera: The UV-Visible camera has a catadioptic telescope          
using fused silica lenses focusing onto a metachrome-coated CCD imager.       
Active wavelength response is limited on the short wavelength end by the      
transmission of fused silica and the optical blur of the lens.                
Wavelength response on the long end is limited by the response of the         
CCD. Six spectral bands can be selected from a filter wheel which is          
controlled through the same serial-addressable synchronous interface          
The Thomson focal plane array (FPA) used is a frame-transfer device,          
accomplishing electronic shuttering by rapidly shifting the active pixel      
area into the storage area, pausing for the (13-bit programmable)             
integration time, then rapidly shifting the captured image into a             
storage buffer from which the image is read out. Post-FPA electronics         
allow three gain states followed by 5 bits of offset that span 248            
counts in the analog regime to augment the basic 8-bit A/D conversion.        
Gain is A/D digitization noise limited, so proper exposure is critical.       
Working against the day side of the Moon as a target, typical                 
integration times were as short as several milliseconds in the lowest         
gain state (1000 electrons/bit) near sub-solar illumination points at         
the brighter spectral bands, increasing to 40 msec near the polar             
regions in the mid-gain setting for the weaker  415 and 1000 nm spectral      
bands. The UV-Visible Performance specifications are shown in Table 2.        
Wide Field of View Star Trackers: The star tracker cameras have a             
concentric optics design with a fiber optic field flattener to couple         
the image surface to a CCD array. The CCD is kept cleared by                  
continuously shifting lines and reading out pixels at the 5 MHz base          
rate, which is approximately 100 microseconds per line. Integration is        
accomplished by stopping this clearing process for the specific time          
(13-bit programmable), then proceeding with a second line-shifting            
operation into a readout buffer line and completing readout at 5 MHz.         
Camera electronics are identical to those of the UV-VIS camera, with the      
gain circuit resistors selected to be slightly more sensitive.                
The primary function of the star tracker cameras was to provide stellar       
images that were processed against an onboard star catalog to find            
pointing, thus establishing absolute angular references for navigation.       
The scientific uses of the cameras were secondary. Only broad band            
operation was available. Owing to the line-transfer electronic                
shuttering, imaging was limited to dim targets such as the lunar surface      
illuminated by earth shine. System noise is about 1.0 bits rms at the         
highest sensitivity setting. The star tracker camera performance              
specifications are provided in Table 3.                                       
High Resolution Imaging (HiRes) camera: The HiRes camera combines a           
lightweight beryllium telescope with an image intensifier-coupled frame       
transfer CCD imager. Image shuttering is accomplished through voltage         
gating of the image intensifier. Maximum integration time is 733              
milliseconds in  10.67 microsecond increments. Spectral response is           
limited in the system by the S-2 photocathode between 0.4 and 0.8             
microns. Five spectral bands can be selected from a filter wheel which        
is controlled through the same SASI interface. A sixth filter position        
is allocated to an opaque filter for the image intensifier's protection.      
The post-FPA electronics circuitry is identical to that used in the UV-       
VIS camera. Images of the day side of the moon used intensifier gate          
times with relatively low gain settings on the order of 1 millisecond.        
Lifetime concerns about the photocathode and micro channel plates in the      
intensifier unit drove operational settings to low exposures. This            
resulted in photon shot noise significantly contributing to the overall       
noise in the HiRes sensor. The HiRes performance specifications are           
provided in Table 4.                                                          
Laser Rangefinder (LIDAR): The LIDAR unit shares the telescope of the         
HiRes camera, splitting the 1064 nm return signal from the NdYag source       
off to an avalanche photodiode (APD) detector with dichroic filter. The       
optics are non-imaging, providing an exit pupil through simple relay          
optics at the APD. The APD electronics includes a temperature                 
compensation feature for the APD bias voltage and programmable                
thresholding of the output signal. The APD current is amplified and           
inverted to a voltage by a transimpedance amplifier with a gain of 230X,      
a low frequency cutoff of 3 MHz, and a high frequency cutoff of 23 MHz.       
The APD current is amplified, then discriminated for changes (increase)       
through 14 MHz discriminator. Voltage changes exceeding the programmed        
threshold are flagged as returns.                                             
Range value is determined by the clock cycles since a laser output            
sampled start pulse is received. The clock counter has only 14 bits           
owing to the hardware availability limitations. In order to allow             
returns up to the 640 km maximum range required in the lunar mission,         
returns from the discriminator are binned 4 to a clock count, turning         
the 23 MHz response into a 40 meter height bin. Internal memory in the        
LIDAR unit saves up to 6 "returns" per laser firing, with up to 4 saved       
between programmable search range minimum/maximum values. Threshold is        
set for the best compromise between missed detection and false alarms.        
The LIDAR components are summarized in Table 5.                               
Near Infrared Camera (NIR): The NIR camera uses a catadioptric lens with      
a 256x256 indium antimonide (InSb) FPA mechanically cooled to cryogenic       
temperature.  The FPA operated at 70 plus-or-minus 0.5 K at the Moon and      
showed excellent stability over the more than 500 hours of operation in       
space.  The lens design features all ZnSe refractive elements with a          
relay to provide an external pupil for 100% efficient cold shielding.         
This lens design was chosen for image quality and focus stability.            
Wavelength range was constrained by the optics and the InSb response to       
somewhat less that 1.0 to 5.5 microns. Six wavebands were selected by         
the NASA advisory committee, all falling well inside this range.              
Camera electronic programmable variables allow 4 integration times, 5         
bits of gain, and 8 bits of offset. Gain states are spaced approximately      
evenly from 0.5 to 36 factors of voltage multiplication. Offset is            
subtracted before gain is applied with 0 V to full well range that can        
be set in 1/255 full well increments. The NIR camera performance              
specifications are shown in Table 6.                                          
Longwave Infrared Camera (LWIR): The LWIR camera uses a catadioptric          
lens with a 128 x 128 mercury cadmium telluride (HCT) FPA. The FPA was        
mechanically cooled to cryogenic temperatures during operation with a         
temperature set point goal of 65 K. The optics design incorporates an         
internal relay with 100% cold shield efficiency. Wavelength ranges was        
controlled by the cold filter to 8.0 to 9.5 microns.                          
Camera electronic design is virtually identical to the NIR camera, with       
a minor alteration in line readout to compensate for the difference in        
column count between the two arrays (256 for the NIR, 128 for the LWIR).      
The NIR and LWIR cameras also share a common cryocooler and dewar             
design, with minor modifications made to accommodate cold shield and          
cold filter differences. The LWIR Performance Specifications are shown        
in Table 7.                                                                   
Table 2  UV/Visible Performance Specifications                                
Focal Plane Array:                                                            
Type              Si Charge Coupled Device                                    
                  Thomson TH7863-CRU-UV                                       
Pixel format      288x384                                                     
Pixel size        23x23 microns                                               
Readout rate      4MHz                                                        
Clear aperture    46nm                                                        
Speed             F/1.96                                                      
Wavelength        0.3 to 1.1 microns                                          
Field of view     4.2 deg. x 5.6 deg.                                         
Pixel IFOV        255 microradians                                            
Point spread      1.1 to 1.5 pixels                                           
Filters           415nm cw (plus-or-minus 20 nm bw)                           
                  750nm (plus-or-minus 5)                                     
                  900nm (plus-or-minus 10)                                    
                  950nm (plus-or-minus  15)                                   
                  1000nm (plus-or-minus 15)                                   
                  400 to 950 broad band                                       
Camera Electronics:                                                           
A/D resolution    8 bits                                                      
Frame rate        10 Hz                                                       
Readout time      27.4 ms                                                     
Integration time  0.2-733 ms                                                  
Digitization gain 150,350,1000 electrons/count                                
Offset control    248 gray levels                                             
Power             4.5 W                                                       
Mass              410 grams                                                   
Envelope          15.5cm x 11.7 cm x 10.4 cm                                  
Common Module Filter Wheel System:                                            
Type              6 position, 90 deg. stepper motor driven,                   
                  Hall effect position sensors                                
Step and Settle time    < 250ms                                               
Position repeatability  < 10mr                                                
Power             0.15 W quiescent, 11.0 W stepping                           
Table 3  Star Tracker Performance Specifications                              
Focal Plane Array:                                                            
Type              Si Charge Coupled Device                                    
                  Thomson TH7863                                              
Pixel format      576x384                                                     
Pixel size        23x23 microns                                               
Readout rate      5MHz                                                        
Clear aperture    14nm                                                        
Speed             F/1.3                                                       
Wavelength        0.5 to 1.1 microns                                          
Field of view     43 deg. x 28 deg.                                           
Pixel IFOV        1.3 microradians                                            
Point spread      2 pixels                                                    
Camera Electronics:                                                           
A/D resolution    8 bits                                                      
Readout time      54.8 ms                                                     
Integration time  0.2-733 ms                                                  
Digitization gain 75, 150, 350 electrons/count                                
Offset control    248 gray levels                                             
Power             4.5 W                                                       
Mass              290 grams                                                   
Envelope          11.7 cm x 11.7 cm x 13.2 cm                                 
Table 4  HiRes Imaging Receiver Performance Specifications                    
Focal Plane Array:                                                            
See UV/Visible specifications in Table 1, except arrays are not UV            
Clear aperture    131 nm                                                      
Speed             F/9.5                                                       
Wavelength        0.4 to 0.8 microns                                          
Field of view     0.4 deg. x 0.3 deg.                                         
Pixel IFOV        18 microradians                                             
Point spread      4 pixels                                                    
Filters           415 nm cw (plus-or-minus 20 nm bw)                          
                  560 nm cw (plus-or-minus 5)                                 
                  650 nm cw (plus-or-minus 5)                                 
                  750 nm cw (plus-or-minus 10)                                
                  400 to 800 nm broad band                                    
Image intensifier:                                                            
Image intensifier module      General Atomics 0131-Z12-2-009                  
Useful photocathode diameter  12 mm                                           
Luminous gain     1000 fL/fC                                                  
Limiting resolution           40 lp/mm                                        
Gain control      8 bits                                                      
Camera Electronics:                                                           
See UV/Visible specifications in Table 1                                      
Power             9.5 W                                                       
Mass              1120 grams                                                  
Envelope          17.0 cm x 18.1 cm x 36.4 cm                                 
Table 5 LIDAR Components                                                      
Type              Si APD C30954E, 0.4 to 1.1 micron                           
Pixel format      single APD                                                  
Pixel size        0.5 mm2                                                     
Pixel FOV         1.0 mrad diameter                                           
APD gain          100 X nominal                                               
Transimpedance amplifier      230 X gain; 3 MHz to 23 MHz response            
Shared path wit HiRes Camera. Dichroic beamsplitter. Exit pupil matched       
to APD. Mass and envelope included in HiRes description.                      
Camera Electronics:                                                           
Range Resolution  40 meter                                                    
Detection         signal derivative                                           
Detection criteria       programmable threshold (8 bits)                      
Laser Transmitter:                                                            
Wavelengths       532 nm/1064 nm                                              
Pulse energy      171 mJ @ 1064, 9 mJ @ 532 nm                                
Pulse width       < 10ns                                                      
Beam divergence   < 500 mrad @ 064 nm, 4 mrad @ 532 nm                        
Shot profile      Continuous @ 1 Hz, 400 shots @ 8 Hz                         
Laser Transmitter Optics:                                                     
Type              5X Galilean-type telescope                                  
Clear Aperture    38 mm exit diameter                                         
Start pulse detector    Analog Modules, Inc. Model 754                        
Pockels cell driver     Analog Modules, Inc. Model 824                        
Laser diode heater      Minco Kapton film resistance type; 16C to 18C         
Power             6.8 W at 1 Hz; 2.6 W quiescent                              
Laser Transmitter Mechanical:                                                 
Mass              635 grams laser head; 615 grams power supply                
Envelope          13.3 cm x 15.2 cm x 3.9 cm high (power supply)              
Table 6  NIR Performance Specifications                                       
Focal Plane Array:                                                            
Type              Amber PV InSb                                               
Pixel format      256x256                                                     
Pixel size        38x38 microns                                               
Non-operable pixels     less than 0.5%                                                
FPA operating temp.     70 K                                                  
FPA well capacity       11.7 million electrons                                
Clear aperture          29nm                                                  
Effective focal length  96 mm                                                 
Cold stop               F/3.33, 6.0 mm diameter                               
Cold shield efficiency  100%                                                  
Field of view     5.6 deg. x 5.6 deg.                                         
Pixel IFOV        400 x 400 microrad                                          
Point spread      greater than 50% energy in 30 micrometer slit                          
Filters           1100 nm (plus-or-minus 30 nm)                               
                  1250 nm (plus-or-minus 30 nm)                               
                  1500 nm (plus-or-minus 30 nm)                               
                  2000 nm (plus-or-minus 30 nm)                               
                  2600 nm (plus-or-minus 30 nm)                               
                  2690 nm (plus-or-minus 60 nm)                               
Camera Electronics:                                                           
A/D resolution    8 bits                                                      
Frame rate        7.1 Hz (single frame mode)                                  
Integration times 11, 33, 57, and 95 ms                                       
Digitization gain 0.5 to 36 X voltage multiplication                          
Offset control    8 bits                                                      
Power             13.0 W                                                      
Type              Ricor K506B integral Stirling with                          
                  H-10 FPA temperature closed-loop                            
                  control electronics                                         
Avg. power        11.0 W steady-state                                         
Mass              1920 grams                                                  
Envelope          10.4 cm x 11.5 cm x 36.5 cm long                            
Table 7  LWIR Performance Specifications                                      
Focal Plane Array:                                                            
Type              Amber PV HgCdTe                                             
Pixel format      128x128                                                     
Pixel size        50x50 microns                                               
Non-operable pixels     < 5%                                                  
FPA operating temp.     65 K nominal                                          
FPA well capacity       42 million electrons                                  
Equivalent clear aperture     29nm                                            
Effective focal length        96 mm                                           
Cold stop               F/2.67, 7.47 mm diameter                              
Cold shield efficiency  100%                                                  
Field of view     1 deg. x 1 deg.                                             
Pixel IFOV        143 x 143 microrad                                          
Point spread      > 60% energy in 79 micrometer slit                          
Camera Electronics:                                                           
A/D resolution    8 bits                                                      
Frame rate        52.9 Hz (single frame mode)                                 
Pixel rate        500kHz                                                      
Integration times 0.115, 0.92, 2.30, and 4.60 ms                              
Digitization gain 0.5 to 36 X voltage multiplication                          
Offset control    8 bits                                                      
Power             13.0 W                                                      
See NIR cryocooler specifications in Table 5.                                 
Mass              2100 grams                                                  
Envelope          14.7 cm diameter x 36.1 cm long                             
Preflight Calibration: Extensive pre-flight calibration data were             
acquired using an automated calibration facility at LLNL. In a typical        
calibration configuration, a sensor was mounted inside an environmental       
chamber whose temperature was set from -20 to 20 deg. C which were the        
expected operating temperatures for the mission. Depending on the             
measurement types, the sensors saw either a flat diffused light source        
or an off-axis collimator with various pinholes as the point source. A        
custom board controlled the sensor parameters from the host computers;        
the video signal was acquired using a commercial image processor. During      
data acquisition many thermal parameters such as FPA and chamber              
temperatures were monitored  and recorded as part of the image                
structure. All calibration processes were fully automated enabling rapid      
data acquisition and minimization of operator error. Pre-flight               
calibration attempted to cover similar light levels expected from the         
lunar surface and spanning the same camera settings required for the          
lunar mapping phase.                                                          
The pre-flight calibration measurements included radiometric                  
sensitivity; FPA uniformity; gain and offset scale factors;                   
temporal/spatial noise; dark noise dependence on FPA temperatures,            
integration times or input voltage levels, spectral response of FPA;          
optical distortion map; point spread function; electronic warm-up time        
and cryocooler cool down time. For the thermally sensitive sensors such       
as LWIR camera, the noise measurement was performed using a vacuum            
chamber to simulate the space thermal environment.                            
Many pre-flight calibration coefficients were applied to lunar data           
showing reasonable agreement with expected performance. In-flight             
calibration data will allow minor corrections for vacuum flight               
condition and sensor degradation over mission lifetime to be added to         
the pre-flight calibration results. The final calibration is expected to      
be better than 5%.                                                            
Data Compression: Data compression was done onboard using the CNES            
compression chip. The processing was performed on a completed, framed         
image prior to storage on the solid state data recorder (SSDR) when the       
appropriate compression flag is set.                                          
The compression chip developed by MATRA under CNES specifications is          
used in two modes, which could be selected via a software uplink              
command. The first mode optimized rms error for a nominal compression.        
The second (JPEG) provided visual optimization at a fixed compression         
rate. In the first mode, blocks of 8x8 pixel 8-bit data are transformed       
to a best fit cosine series expansion in the orthogonal row and column        
directions. This algorithm tends to preserve high frequency information       
with less data loss than does JPEG at the same compression ratio for the      
lunar data. Total signal from the 8x8 block is preserved exactly. The         
nominal amount of compression was set by limiting the scene error             
induced by compression to a fraction of the camera's temporal noises.         
Analysis of lunar images during the first part of the mission showed          
that the quantization matrix used by the chip was optimum for the             
imaging cameras. The HiRes camera, however, was operated in JPEG mode.        
The high frequency information in the HiRes scenes was spurious (it was       
caused by gain non-uniformity of the intensifier tube); eliminating high      
frequency content allowed higher compression without harming the              
information content of the scenes. The average compression rate for all       
images obtained during the mission was 5.5.                                   
5. Lunar Orbital Dynamics                                                     
Based on the characteristics of the baseline sensor complement, the           
mapping of 100% of the lunar surface was done in two lunar days (two          
Earth months).  During the nominal two month mapping mission, the             
required image overlap for the UV/VIS and NIR cameras was ~15% in the         
down track and ~10% in the cross track directions.  This required that        
the periselene of the lunar orbit be maintained at an altitude of 425         
plus-or-minus 25 km.  In order to image 100% percent of the moon's            
surface during the two months, the spacecraft was required to be in a         
polar orbit.  This requirement was satisfied with the inclination of the      
orbit at 90 degrees plus-or-minus 1 degree with reference to the lunar        
equator.  To provide the necessary separation for the alternating             
imaging strips to cover the entire surface of the moon during the two         
months, the orbital period was approximately 5 hours.  During this            
orbital period the moon rotated approximately 2.7 degrees beneath the         
spacecraft.  The orbit also was a sufficiently long period to allow the       
transmission to Earth of data collected during the imaging phase of each      
During the lunar mapping phase of the mission, there were four separate       
observational periods. The first, orbits 1-31, was a shakedown and            
testing period where the spacecraft observation sequences were tested         
and refined. Observations of special targets, such as Apollo landing          
sites, were additionally acquired during this period. The second period,      
orbits 32-168, was the first month's systematic mapping with periselene       
in the southern hemisphere. The third period, orbits 169-297, was the         
second month's systematic mapping with periselene at the northern             
hemisphere. In the fourth period, orbits 298-348, periselene remained in      
the northern hemisphere. In this period, observations were made to cover      
gaps in coverage, acquire observations of special targets, acquire            
stereo observations over Orientale Basin, and obtain calibration data.        
Clementine left lunar orbit soon after orbit 348.                             
The best data for lunar mineral mapping is obtained if the  solar phase       
angle is less than 30 degrees.  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 degrees, the plane of the lunar      
orbit should contain the Moon-Sun line half way through the two-month         
lunar mapping period.  Therefore, insertion into the lunar orbit was          
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.       
The angle between the Moon-Sun line and the orbital plane was close for       
approximately five weeks before becoming zero. Table 8 contains a list        
of Clementine's orbital parameters                                            
            Table 8 Clementine Orbital Parameters                             
Orbital Period:         4.970 hr < P < 5.003 hr                               
Radius of Periselene:   2138 km < radius < 2188 km                            
Eccentricity:           0.35821 < e < 0.37567                                 
Right Ascension:        -3 deg < Omega < +3 deg (J2000)                        
Inclination:            89 deg < i < 91 deg                                   
Periselene:            -28.4 deg < w < -27.9 deg (1st month)                  
                        29.6 deg < w <  29.2 deg (2nd month)                  
Orbit determination and monitoring was done on a continuous basis             
throughout the lunar pre-mapping phase.  The gravitational potential          
field of the moon has not been fully mapped, and large lunar mass             
concentrations may have had a significant perturbation effect on the          
orbit.  Maintenance burns were required to maintain the orbit within the      
required envelope.  The number of these burns was minimized to avoid          
unnecessary disruptions to the systematic mapping.  To this end, any          
required periapsis burns were performed away from periselene in the           
direction of the near pole.                                                   
Attitude measurement accuracy was necessary to determine spacecraft           
pointing to within 0.03 degree, 0.5 milliradians. This accuracy was           
achieved in real-time, in darkness or sunlight throughout the lunar           
mapping phase.  The spacecraft was three-axis stabilized and capable of       
autonomous, open loop inertial pointing with an accuracy of 0.05 degree,      
0.87 milliradian, or better.  This accuracy was required to support use       
of the high resolution camera because of its narrow field of view for         
imaging selected target sites during the lunar mapping mission.               
The spacecraft was able to execute controlled, relative pointing motion       
about a pointing vector for scanning across targets.  The relative            
pointing was capable of controlled motion of 75 microradians.                 
During lunar imaging, the spacecraft had to maintain a NADIR pointing         
attitude.  This required a greater than 180 degree rotation over the          
approximately 1.5-2.0 hour period during each lunar orbit.  The               
spacecraft was also required to maintain an angular bias about the X-         
axis from NADIR to permit an imaging groundtrack parallel but offset          
from the NADIR groundtrack.                                                   
The spacecraft was required to point to the Earth center, and to a            
specified tracking station site on the Earth, for the dumping of data         
using the high-gain directional antenna.                                      
To help accomplish attitude determination, the spacecraft had two             
inertial measurement units (IMU) and two star trackers.  Because of a         
solar exclusion angle constraint, one of the two star trackers had to be      
covered during lunar orbit.  To meet the aforementioned pointing              
requirements, during lunar orbit a star tracker image was processed and       
the spacecraft attitude knowledge was updated at 10 second intervals or       
6. Initial Scientific Results                                                 
Over the course of 71 days in lunar orbit, Clementine systematically          
mapped the 38 million square kilometers of the Moon in eleven colors in       
the visible and near infrared parts of the spectrum (nearly 1,000,000         
images). In addition, the spacecraft took 620,000 high resolution and         
about 320,000 mid-infrared thermal images, mapped the topography of the       
moon with a laser ranging experiment, improved our knowledge of the           
surface gravity field of the Moon through radio tracking, and carried a       
charged particle telescope to characterize the solar and magnetospheric       
energetic particle environment. All sensors on the spacecraft met or          
exceeded expectations of their performance. The first global color view       
of the Moon was acquired, major compositional provinces were identified,      
and geology and composition details were mapped.                              
The images from Clementine constitute the first color global digital          
data set of the Moon. The NASA Science Team advised the project on the        
selection of color filters for the two principal mapping cameras: the         
UV/Visible camera and the NIR camera. The color of the Moon in the            
visible to near-infrared part of the spectrum is sensitive to variations      
in both the composition of the surface material and the amount of time        
material has been exposed to space. The Clementine filters were selected      
to characterize the broad lunar continuum and to sample parts of the          
spectrum that are known to contain absorption bands diagnostic of iron-       
bearing minerals and plagioclase feldspar, the dominant mineral               
constituents of the lunar crust. By combining information obtained            
through several filters, multispectral image data are being used to map       
the distribution of rock and soil types on the Moon.                          
Clementine was successful in systematically mapping the Moon in these 11      
colors at an average surface resolution of about 200 meters per pixel.        
The initial examination of the data attests to its excellent quality.