PDS_VERSION_ID = PDS3 RECORD_TYPE = STREAM LABEL_REVISION_NOTE = "20090629, L. Gaddis - Initial Version 20101001 C. Isbell - Post Review" OBJECT = INSTRUMENT_HOST INSTRUMENT_HOST_ID = "LO3" OBJECT = INSTRUMENT_HOST_INFORMATION INSTRUMENT_HOST_NAME = "LUNAR ORBITER III" INSTRUMENT_HOST_TYPE = "SPACECRAFT" INSTRUMENT_HOST_DESC = " LUNAR ORBITER III MISSION OVERVIEW ================================== Much of the information in this document was abstracted from the Lunar Orbiter mission pages of the the National Space Science Data Center Web site (http://nssdc.gsfc.nasa.gov/nmc/masterCatalog.do?sc=1967-008A). See references cited for more detail. Initiated in early 1964, the NASA Lunar Orbiter program included the design, development, and utilization of complex automated spacecraft technology to support the acquisition of detailed photographs of the lunar surface from orbit to support the Apollo Program [e.g., Boeing Company, 1967, 1968a, b; Hansen, 1970; Bowker and Hughes, 1971; Byers, 1977]. Management of all five of the Lunar Orbiter missions was assigned to NASA Langley Research Center. The Boeing Company served as the prime spacecraft contractor. Eastman Kodak Company was responsible for the photographic system. For all Lunar Orbiter experiments, data were collected by a suite of 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. DSN consisted of the Space Flight Operations Facility (SFOF) and Deep Space Stations (DSS) to provide two-way communications with the spacecraft, data collection and data processing. Goddard Space Flight Center was responsible for the worldwide network of communication lines between the tracking stations and SFOF. The following sections provide an overview of the Lunar Orbiter III spacecraft, the various subsystems it carried, and the DSN ground system. The Lunar Orbiter III mission obtained photographs of areas of the lunar surface for confirmation of safe landing sites for the Surveyor and Apollo missions. The primary objective of Lunar Orbiter III was to re-photograph the most promising lunar landing sites photographed by Lunar Orbiter I and II and to help determine landing vehicle configuration and verify vehicle design. The Mission also collected selenodetic, radiation intensity, and micrometeoroid impact data. LUNAR ORBITER III SPACECRAFT SUMMARY ==================================== The Lunar Orbiter III spacecraft was a three-axis stabilized vehicle with a normal mass of 383 kg and was designed to be mounted within an aerodynamic nose shroud on top of the Atlas/Agena launch vehicle [Boeing Company, 1968b]. The main bus had the general shape of a truncated cone, 1.65 meters tall and 1.5 m in diameter at the base. The spacecraft was composed of three decks supported by trusses and an arch. The equipment deck at the base of the craft held the battery, transponder, flight programmer, inertial reference unit (IRU), Canopus star tracker, command decoder, multiplex encoder, traveling wave tube amplifier (TWTA), and the photographic system. Four solar panels were mounted to extend out from this deck with a total span across of 3.72 m. Solar panels were approximately perpendicular to the spacecraft centerline. Power of 375 W was provided by four solar arrays containing 10,856 n/p solar cells that ran the spacecraft and charged the 12 A-h nickel-cadmium batteries. The batteries were used during brief periods of occultation when no solar power was available. In the deployed configuration, extending out from the base of the spacecraft were a 1 meter diameter high-gain antenna on a 1.32 m boom for transmission of photographs and an omnidirectional low-gain antenna on a 2.08 m boom for other communications. A 10 W transmitter relayed photographs and a 0.5 W transmitter was used for other communications. Both antennas operated in S- band at 2295 Hz. Propulsion for major maneuvers was provided by the gimballed velocity control engine, a hypergolic 100 pound-force (445 N) thrust Marquardt rocket motor. Above the equipment deck, the middle deck held the velocity control engine, propellant, oxidizer and pressurization tanks, Sun sensors, and micrometeoroid detectors. The micrometeoroid detectors were located in a ring just below the fuel tanks. These devices were pressurized cans of known thickness that, when punctured, lost pressure and sent a signal of this event to Earth. Two proton-radiation detectors were carried for mission control purposes. Shielding these detectors approximates that at two critical locations within the photographic system, and telemetered dose values and rates make possible an evaluation of the gravity of any solar proton activity encountered. The third deck consisted of a heat shield of multilayered aluminized Mylar and Dacron to protect the spacecraft from the firing of the velocity control engine. The nozzle of the engine protruded through the center of the shield. Mounted on the perimeter of the top deck were four attitude control thrusters. LAUNCH VEHICLE ============== The first stage of the launch vehicle was an SLV-3 (Atlas) [e.g., Boeing Company, 1968b]. The vehicle placed the upper stage into proper coast ellipse; relayed commands for separation of the upper stage vehicle and started the Agena primary timer; and relayed commands to the Atlas-Agena interface to jettison the shroud and start the secondary timer of the launch vehicle. The secondary objective was to use telemetry data to measure Atlas performance. The second stage of the launch vehicle was an Agena-D. The Agena placed the spacecraft into a lunar-coincident transfer (cislunar) trajectory within prescribed orbit dispersions, and performed Agena attitude and retro- maneuvers following Agena-spacecraft separation to ensure that the Agena would not intercept the spacecraft, pass within 20 degrees of the center of the Canopus tracker field of view, or impact the Moon. The secondary objective was to provide tracking and telemetry data for the evaluation of Agena performance. All objectives of the Lunar Orbiter III launch were satisfied. Atlas-Agena separation was properly accomplished and good telemetry data was obtained for Atlas systems analysis. Agena arrived in the lunar vicinity approximately 6 hours after the spacecraft and was approximately 17,000 km beyond lunar capture. SPACECRAFT SUBSYSTEMS ===================== Propulsion Subsystem -------------------- Propulsion for major maneuvers was provided by the gimballed velocity control engine, a hypergolic 100 pound-force (445 N) thrust Marquardt rocket motor [e.g., Boeing Company, 1967]. The motor consists of two oxidizer tanks, two fuel tanks, and a liquid propellant rocket engine. The propellants utilized are nitrogen tetroxide and Aerozine-50. During thrust, the velocity change, as detected by a spacecraft accelerometer, was compared with requirements stored in the flight programmer to determine thrust duration. When the flight programmer commanded the rocket engine valves to open, gas pressure, acting on propellant tank bladders, forced the fuel and oxidizer into the engine. No ignition system was required because the propellants ignited when mixed, and thrust continued until the engine valves were closed. Although the spacecraft trajectory was established by the ATLAS/AGENA launch vehicle, minor changes to the translunar trajectory and the velocity changes required for orbital insertion was accomplished by the spacecraft. An altitude maneuver established the direction of thrust. Operation and performance of the subsystem was well within specification throughout the mission. Three propulsive maneuvers were conducted in support of the primary mission; these were: 5.09-mps midcourse, 704.3-mps orbit injection, and 50.7-mps orbit transfer. Attitude Control Subsystem -------------------------- The Attitude Control Subsystem controls the execution of all spacecraft events and maneuvers for example to precisely position the spacecraft for picture taking, velocity changes or orbit transfers [e.g., Boeing Company, 1967]. The subsystem provided both three-axis stabilization and attitude control. These were provided by four one-lb nitrogen gas jets. The attitude reference for yaw and pitch was provided by a Sun sensor in the equipment mounting plate so that the solar panels normally faced the Sun. The roll axis reference was provided by an electro-optical sensor that tracked the star Canopus. This attitude resulted in the high-gain antenna being pointed toward Earth, with an assist from a rotatable boom on the unit. Signals from these sensing devices controlled N2 gas ejection from attitude control jets to acquire and maintain the necessary spacecraft orientation. Stable thrust vector control of the spacecraft attitude was maintained through three velocity control engine burns. The spacecraft departed from the Sun-Canopus orientation only to point the liquid rocket for velocity changes or to point its camera for photography. Gyros in the inertial reference unit maintain attitude whenever the Sun or Canopus is hidden. The attitude control subsystem maintained spacecraft orientation with respect to the Sun and Canopus on command within +-0.2 and 2.0 degrees. Attitude control was maintained with the spacecraft pitched from 15 to 45 degrees away from the Sun for approximately 56% of the mission. Throughout all phases of the mission, the spacecraft attitude was accurately controlled, receiving commands through the flight programmer to perform 383 single-axis maneuvers. Maneuver accuracy of the subsystem was within the design tolerance. The subsystem satisfied all mission objectives. Power Subsystem --------------- Power of 375 W was provided by the four solar arrays containing 10,856 n/p solar cells which would directly run the spacecraft and also charge the 12 Aúh nickel-cadmium battery [e.g., Boeing Company, 1967]. Solar panels provided all necessary electrical power during the spacecraft maneuvers or occultation of the Sun. Electrical power was supplied by a 28-volt rechargeable nickel-cadmium storage batteries. Power regulators and controllers protected the solar panels, battery, and spacecraft subsystems from unusual power fluctuations. The number of solar cells provided allowed for the possibility that some cells might fail or be damaged by micrometeorites during the mission. Power subsystem performance during Mission III was completely satisfactory. No constraints were imposed on flight operations beyond the requirement that array illumination be sufficient to meet the demands imposed by spacecraft electrical loads and energy balance requirements. The batteries were used during brief periods of occultation when no solar power was available. Thermal Control Subsystem ------------------------- The thermal control subsystem of the Lunar Orbiter spacecraft is a passive system mounted on a Sun-oriented equipment mounting deck (EMD) [e.g., Boeing Company, 1967]. Thermal control was maintained by a multilayer aluminized Mylar and Dacron thermal blanket which enshrouded the main bus, special paint, insulation, and small heaters. Heat generated by equipment was conducted to the EMD, where it was radiated to the space environment. The EMD was coated with a low-solar-absorbance paint. Thermal control was achieved by varying the attitude of the EMD with respect to the Sun. Supplemental heating was supplied, as needed, to the propellant tanks and photo subsystem by electric heaters. Spacecraft temperatures were maintained within prescribed temperature limits throughout the mission - with the exception of Orbit 149, during which the film-drive motor failure occurred. During this period, from 061:16:00 to 061:18:00 GMT, the temperatures of PT01 and PT02 increased to a maximum of 88.8 and 93.9øF, respectively. The normal temperature levels for these channels during preceding orbits, with approximately the same spacecraft attitude {31 to 34 degrees off Sun) were in the 73 to 7SøF range. Photographic Subsystem ---------------------- The photographic subsystem was designed to photograph the lunar surface, process the exposed, film, scale the processed film with a flying-spot scanner, and provide video signals to the communications subsystem for transmission to Earth [e.g., Beeler and Michlovitz, 1969; Boeing Company, 1968a, b; Hansen, 1970; Kosofsky and El-Baz, 1970; Anderson and Miller, 1971]. The self-contained, dual-camera system was enclosed in a pressurized container to maintain the environmental conditions required for operation of a photographic laboratory in a space environment. The Mission III photo subsystem was equipped with a 0.21 neutral-density filter in front of the 80- mm lens. The resultant 80-mm lens transitivity was 59%; the 610-mm lens transmissivity was 65%. The two cameras operated simultaneously, placing two discrete film exposures on a common roll of 70-mm film. Each camera operated at a fixed f/5.6 lens aperture, at shutter speeds of either 1/25, 1/50, or 1/100 second. The high- resolution (H) frame was exposed through a 610-mm narrow-angle lens and a focal plane shutter. The medium-resolution (M) frame was exposed through an 80-mm wide-angle lens and a between-the-lens shutter. This lens provided angular coverage of 44.4o by 38o. The 610 mm focal length lens photographed a small area centered on this field with an angular coverage of 20.4o by 5.16o. Photo subsystem performance was generally satisfactory from launch through the Orbit 149 anomaly. Film handling was as not as good as in preceding missions and the camera film advance showed some small effects of film set. The hang-ups, which occurred throughout the mission, may have reduced the total amount of data retrieved from the spacecraft, but had no effect on photo quality. Only about 75% of the photographs were transmitted before a failure in the film-advance motor terminated the readout operation. After photo subsystem status became apparent, considerable effort was devoted to further testing and to maximizing the readout looper contents. The Lunar Orbiter III photographic systems performance was evaluated in terms of resolving power in lines per millimeter resolved in the final image under specified operating conditions, and in modulation transfer, or percentage of the object plane information content transferred to the final image plane. Eastman Kodak Company specified 76 lines per millimeter at 3:1 contrast from a lunar altitude of 46 kilometers. The prime contractor accepted signal-to- noise ratio in the final ground film of at least 3:1 when photographing a cone « meter high with a base diameter of 2 meters under typical lunar lighting conditions. By both measures, the photographic system exceeded predictions. Communications Subsystem ------------------------ The Lunar Orbiter communications subsystem served to transmit telemetry and video data to Earth and receive spacecraft commands from Earth [e.g., Bundick et al., 1965]. Communications were via a 10 W transmitter and the directional 1 m diameter high-gain antenna for transmission of photographs and a 0.5 W transmitter and omnidirectional low-gain antenna for other communications. Both antennas operated in S-band at 2295 MHz. High- and low-gain antennas performed satisfactorily throughout the mission. The antennas were deployed after Agena separation. Most functions aboard the spacecraft were controlled by the onboard programmer. This united received commands from Earth stations and executed them immediately or stored them for execution at a precise later time. Sufficient memory slots were available to automatically control spacecraft functions for periods of several hours. The communications subsystem also provided Doppler and ranging signals used by the DSIF for tracking purposes. All incoming signals were received by the low-gain antenna. A transponder responded to the RF carrier and range code to assist the Deep Space Stations in obtaining Doppler tracking and range data. Commands from the DSS were routed to the command decoder and stored. The command, as received was transmitted to Earth where it was checked for accuracy. If verified, an execute command was transmitted to the spacecraft and information stored in the decoder was advanced to the flight programmer. The subsystem performed satisfactorily throughout the mission. All photo data presented to the communications subsystem was successfully processed and transmitted by the spacecraft through the mission and at missions' completion, all subsystem components were functioning satisfactorily. Transponder performance was satisfactory even though two anomalies were observed. Telemetry Subsystem ------------------- The Air Force Eastern Test Range (AFETR), Deep Space Network (DSN) and Manned Space Flight Network (MSFN) were the elements of the Tracking and Data System (TDS) that together supported the tracking and telemetry requirements for the Lunar Orbiter III launch [e.g., Boeing Company, 1968b]. Tracking during the launch phase consisted of C-band tracking of the launch vehicle and reception of VHF and S-band telemetry from the launch vehicle and spacecraft, respectively. The ability to satisfy the near-Earth phase tracking and telemetry requirements was dependent on trajectory characteristics and TDS facilities during that phase. The cislunar injection point for the launch was on an azimuth of 81.6 degrees near the western edge of Africa in the Atlantic Ocean. Initial receipt of telemetry data via the DSN primary sites occurred 66 minutes after liftoff. DSN 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. DSN provides 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, CA), Robledo (near Madrid, Spain), and Tidbinbilla (near Canberra, Australia). Performance of the ground equipment at the Deep Space Stations was satisfactory. The Space Flight Operations Facility (SFOF) provided the mission control center and the facilities to process and display data to support operational mission control. Facilities were provided for the ground reconstruction equipment and for analysis of the reconstructed lunar photographs; there were also facilities for reproduction and distribution of operational data and for microfilming all computer program output. The performance of the entire data system at the SFOF was satisfactory. The telemetry processing station (TPS) and internal communications system at the SFOF provided tracking and telemetry data from teletype and high-speed data line to the SFOF computers and teletype data to the operations areas. The computer complex provided telemetry data processing, tracking data processing, command generation, and command verification. The central computer complex consisted of three computer strings. The entire system performed exceptionally well." 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