| INSTRUMENT_HOST_DESC |
Instrument Host Overview
========================
For most Clementine experiments, data were collected by
instruments on the spacecraft. Those data were then relayed
via the telemetry system to stations of the NASA Deep Space
Network (DSN) on the ground. Radio Science experiments (such
as radio tracking of the spacecraft and bistatic radio
scattering experiments) required that DSN hardware also
participate in data acquisition. The following sections
provide an overview first of the spacecraft and then of the
DSN ground system as both supported Clementine science
activities. The Naval Research Laboratory (NRL) also
operated a tracking station at Pomonkey, Maryland. Air Force
Satellite Control Network (AFSCN) Remote Tracking Stations
(RTS); a station of the Centre Nationales d'Etudes Spatiales
(CNES) near Pretoria, South Africa; stations of the Navy Space
Surveillance Network (NAVSPASUR); and a NASA site in Chile also
participated in radio commanding and/or tracking. But these
were incidental to acquisition of science data and are not
described further.
Instrument Host Overview - Spacecraft
=====================================
The Clementine spacecraft was built at the US Naval Research
Laboratory in Washington, DC. It carried sensors, attitude
control systems and software designed and built by the
Lawrence Livermore National Laboratory (LLNL). The United
States Air Force (USAF) supplied advanced lightweight
composite structures and the launch vehicle, a Titan IIG
refurbished ICBM. Low spacecraft mass was achieved by
incorporating many lightweight technologies developed
through the research and development activities of the
Strategic Defense Initiative (SDI). NASA provided
communications support through the Jet Propulsion
Laboratory's (JPL) Deep Space Network; orbit determination
and operations support were provided from both the NASA
Goddard Space Flight Center and JPL. Supporting these
laboratories were scores of industrial contractors.
The spacecraft consisted of an octagonal prism about 2
meters high. The all-aluminum frame was of conventional
construction; the mid-deck was an aluminum honeycomb.
Spacecraft dry mass was about 230 kg; an approximately
equal mass of liquid fuel was added. Total mass in the
launch configuration was 1690 kg, with most of the weight
in the solid rocket motor (SRM) required for translunar
insertion.
The main instrumentation on Clementine consisted of four
cameras, one with a laser-ranging system. The cameras
included an ultraviolet-visible (UVVIS) camera, a long-
wavelength infrared (LWIR) camera, the laser-ranger (LIDAR)
high-resolution (HIRES) camera, and a near-infrared (NIR)
camera. In addition the spacecraft had two star tracker
cameras (A-STAR, B-STAR), used mainly for attitude
determination; these also served as wide-field cameras for
various scientific and operational purposes. The sensor
package had a mass of 8 kg. The sensors were all located on
one side of the spacecraft body, 90 degrees from the solar
panels. The spacecraft also carried a charged particle
telescope to characterize the spacecraft environment.
Radio tracking data from the spacecraft S-band transponder
provided information on the lunar gravity field. A small
set of bistatic radar experiments was conducted using the
radio transmitter to determine the scattering properties of
the lunar surface.
Propulsion Subsystem
--------------------
The Propulsion Subsystem provided three functions: (1) firing
the SRM for translunar injection, (2) providing three-axis
and spin-stabilized attitude control, and (3) providing
trajectory changes and trims after separation of the SRM.
The STAR 37FM SRM provided 3115 m/sec delta-V capability
for translunar injection. Its thrust specification was
47260 N (nominal) and 54799 N (maximum). The SRM mass was
1152 kg. For the SRM burn, the spacecraft was spun up to
60 rpm using 4.4 N thrusters, then despun using opposing
4.4 N thrusters following the burn. The SRM was separated
from the remainder of the spacecraft after translunar
injection.
A single 489 N bi-propellant thruster mounted on one end
of the prism provided 1744 m/sec delta-V capability after
the SRM separation. It was used to adjust the phasing orbit,
for lunar orbit insertion (550 m/sec), for lunar orbit
trims (including the change in periselene latitude), and
for departure from lunar orbit (540 m/sec). It was fueled
by nitrogen tetraoxide and monomethyl hydrazine (N2O4/MMH);
195 kg of fuel was stored in four tanks; pressurization
was provided by 1 kg of helium stored in two tanks.
Ten fine (4.4 N) and two coarse (22.2 N) monopropellant
thrusters provided attitude control for slewing during
lunar mapping, for spin-up and spin-down, for momentum
dumping, and for active nutation control. These thrusters
were fueled by 55.5 kg of hydrazine in a single tank.
Attitude Control Subsystem
--------------------------
The Attitude Control Subsystem provided both three-axis
and spin-stabilized control. It included two Inertial
Measurement Units (IMUs), two Star Tracker Cameras (STCs),
four reaction wheels, and the twelve monopropellant
thrusters described above. The system was designed for
attitude control of better than 0.05 degrees and knowledge
of better than 0.03 degrees.
The Star Tracker Cameras were mounted on opposite sides
of the spacecraft, allowing at least one to obtain star
field coverage at all times without interference from the
Sun. Each could provide three-axis attitude determination
from a single image provided the Sun, Moon, and Earth were
not in the field of view.
The four reaction wheels (2 Nms) provided attitude control
for fine pointing and for low acceleration slews. The
Attitude Control subsystem normally used three orthogonal
reaction wheels for control, keeping the fourth (mounted
at equal angular displacements from each of the other
three) in reserve.
Electric Power Subsystem
------------------------
The Electric Power Subsystem included a pair of
independently gimbaled, single axis, GaAs/Ge solar
arrays providing a total spacecraft power of up to 360
watts at 30 Vdc, with a specific power of 240 w/kg. The
two arrays protruded from opposite sides of the prism; by
rolling the spacecraft and rotating the panels, full
solar illumination of the panels could be achieved.
The solar arrays were used to charge a 15 A-h, 22-cell,
47-w hr/kg, Nihau common pressure vessel battery. Power
control distribution electronics distributed, conditioned,
and monitored use of electrical power. A Sensor Power
Distribution System performed power distribution and
conditioning for imaging sensors.
Thermal Control Subsystem
-------------------------
The Thermal Control Subsystem maintained internal
spacecraft temperatures within design limits. External
spacecraft surfaces were used under heat-generating
units to radiate excess thermal energy. Multi-layered
insulation blankets covered all non-radiating external
surfaces.
A 4.1 kg beryllium block served as a thermal capacitor,
storing excess heat when radiating surfaces were unable
to dissipate it quickly enough. Diode heat pipes carried
heat from the block to radiating surfaces and protected
against reverse flow.
Sixty thermostats and heaters provided active protection
against cold to various boxes, tanks, and propellant lines
throughout the spacecraft.
Command, Telemetry, and Data Handling Subsystem
-----------------------------------------------
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 safe mode,
attitude control system, and housekeeping operations. A
reduced instruction set computer (RISC) 32-bit processor,
capability of 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 was a state-of-the-art
image compression system provided by the French Centre
Nationale d'Etudes Spatiale (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. They were later
transferred to the ground stations using a 128 kb/s S-band
downlink. The spacecraft was commanded from the ground
using a 1 kb/s S-band uplink from either the NASA Deep Space
Network or from a DoD station.
Demonstration of autonomous navigation, including autonomous
orbit determination, was a major goal of the Clementine
mission. Autonomous operations were conducted in lunar
orbit.
Spacecraft Coordinate System
----------------------------
The spacecraft +Z axis vector was in the nominal
direction of the instrument sensor panel. The +X axis
vector was parallel to the nominal direction of the main
thruster nozzle, and the -X axis vector was parallel to
the spacecraft high-gain antenna (HGA) boresight. The
+Y axis vector formed a right-handed coordinate system
and was in the nominal direction of the solar panel
rotation axis. The low-gain 'omni-directional'
antennas were mounted on the +Z and -Z sides of the
spacecraft body. The spacecraft velocity vector was in
approximately the -X direction when the spacecraft was
oriented for nadir viewing. In the figure below the
viewer is facing the sensor panel; the viewer is looking
in the -Z direction.
+X Axis
^
|
|
Main Thruster
\-----/
--------------- \ / --------------
| | | ------------- | | |
| | | | | | | |
| | | | | | | |
+Y | | | | Sensor | | | |
<--- | Solar | Panel |--| |--| Solar | Panel |
Axis | | | | Panel | | | |
| | | | | | | |
| | | | | | | |
| | | ------------- | | |
--------------- / High-Gain \ ---------------
/ Antenna \
---------------
|
|
V
-X Axis
For more information see [REGEONETAL1994].
Instrument Host Overview - DSN
==============================
The Clementine Radio Science investigations utilized
instrumentation with elements both on the spacecraft and at
the NASA Deep Space Network (DSN). Much of this is shared
equipment, being used for routine telecommunications as
well as for Radio Science.
The Deep Space Network is a telecommunications facility managed
by the Jet Propulsion Laboratory of the California Institute of
Technology for the U.S. National Aeronautics and Space
Administration.
The primary function of the DSN is to provide two-way
communications between the Earth and spacecraft exploring the
solar system. To carry out this function the DSN is equipped
with high-power transmitters, low-noise amplifiers and receivers,
and appropriate monitoring and control systems.
The DSN consists of three complexes situated at approximately
equally spaced longitudinal intervals around the globe at
Goldstone (near Barstow, California), Robledo (near Madrid,
Spain), and Tidbinbilla (near Canberra, Australia). Two of the
complexes are located in the northern hemisphere while the third
is in the southern hemisphere.
The network comprises four subnets, each of which includes one
antenna at each complex. The four subnets are defined according
to the properties of their respective antennas: 70-m diameter,
standard 34-m diameter, high-efficiency 34-m diameter, and 26-m
diameter.
These DSN complexes, in conjunction with telecommunications
subsystems onboard planetary spacecraft, constitute the major
elements of instrumentation for radio science investigations.
For more information see [ASMAR&RENZETTI1993].
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