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
APPENDIX - ORBIT TIME LINE
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
small.
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
volumes.
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
observations.
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
(SASI).
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
Optics:
Clear aperture 46nm
Speed F/1.96
Imaging:
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
Mechanical:
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
Optics:
Clear aperture 14nm
Speed F/1.3
Imaging:
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
Mechanical:
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
enhanced.
Optics:
Clear aperture 131 nm
Speed F/9.5
Imaging:
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
Mechanical:
Mass 1120 grams
Envelope 17.0 cm x 18.1 cm x 36.4 cm
Table 5 LIDAR Components
Detector:
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
Optics:
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
Optics:
Clear aperture 29nm
Effective focal length 96 mm
Cold stop F/3.33, 6.0 mm diameter
Cold shield efficiency 100%
Imaging:
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
Cryocooler:
Type Ricor K506B integral Stirling with
H-10 FPA temperature closed-loop
control electronics
Avg. power 11.0 W steady-state
Mechanical:
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
Optics:
Equivalent clear aperture 29nm
Effective focal length 96 mm
Cold stop F/2.67, 7.47 mm diameter
Cold shield efficiency 100%
Imaging:
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
Cryocooler:
See NIR cryocooler specifications in Table 5.
Mechanical:
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
orbit.
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
less.
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.