PDS_VERSION_ID = PDS3 LABEL_REVISION_NOTE = " 1995-01-01 UNK estimated initial release date; 2009-07-29 NAIF:Semenov added missing RECORD_TYPE keyword; " RECORD_TYPE = STREAM OBJECT = INSTRUMENT_HOST INSTRUMENT_HOST_ID = "CLEM1" OBJECT = INSTRUMENT_HOST_INFORMATION INSTRUMENT_HOST_NAME = "CLEMENTINE 1" INSTRUMENT_HOST_TYPE = "SPACECRAFT" 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]. " END_OBJECT = INSTRUMENT_HOST_INFORMATION OBJECT = INSTRUMENT_HOST_REFERENCE_INFO REFERENCE_KEY_ID = "ASMAR&RENZETTI1993" END_OBJECT = INSTRUMENT_HOST_REFERENCE_INFO OBJECT = INSTRUMENT_HOST_REFERENCE_INFO REFERENCE_KEY_ID = "REGEONETAL1994" END_OBJECT = INSTRUMENT_HOST_REFERENCE_INFO END_OBJECT = INSTRUMENT_HOST END