PDS_VERSION_ID = PDS3 RECORD_TYPE = STREAM LABEL_REVISION_NOTE = "2009629, L. Gaddis - Initial Version 20101001 C. Isbell - Post Review" OBJECT = INSTRUMENT_HOST INSTRUMENT_HOST_ID = "LO5" OBJECT = INSTRUMENT_HOST_INFORMATION INSTRUMENT_HOST_NAME = "LUNAR ORBITER V" INSTRUMENT_HOST_TYPE = "SPACECRAFT" INSTRUMENT_HOST_DESC = " LUNAR ORBITER V 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/spacecraftDisplay.do?id=1967-075A). 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 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 spacecraft, the various subsystems it carried, and the DSN ground system. Lunar Orbiter V completed the tasks of the previous missions with near vertical, convergent telephoto stereo and oblique images of proposed Apollo landing sites. In addition, Mission V obtained close up photographs of geologically interesting features on the Moon's nearside and broad survey images of the previously unphotographed parts of the Moon's far side. LO V also collected selenodetic, radiation intensity, and micrometeoroid impact data and was used to evaluate the Manned Space Flight Network tracking stations and Apollo Orbit Determination program [e.g., Boeing Company, 1968a]. LUNAR ORBITER V SPACECRAFT SUMMARY ================================== The Lunar Orbiter 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, 1968a, b]. 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 battery. 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. The spacecraft performed satisfactorily during the 90-hour cislunar trajectory and 145 lunar orbits, including the multi-axis maneuvers and maneuvers required for photography and to transmit data. The lack of any spacecraft irregularity during the photo mission contributed to the satisfactory completion of the mission. Battery power was rarely required during this mission because the high orbit kept the spacecraft from experiencing Earth occultation periods. 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 launch were satisfied. Atlas-Agena separation was properly accomplished and good telemetry data was obtained for Atlas systems analysis. SPACECRAFT SUBSYSTEMS ===================== Velocity Control Subsystem -------------------------- The velocity control subsystem provides the velocity change capability required for mid=course correction, lunar orbit injection, and orbit adjustment as required. The spacecraft includes a 100-pound-thrust, gimbaled, liquid-fuel rocket engine. The propulsion system uses a radiation-cooled biopropellant liquid rocket engine that employs nitrogen tetroxide as the oxidizer and Aerozine-50 by weight of hydrazine and unsymmetrical dimethylhydrazine as the fuel. The propellants are expelled from the tanks by pressurized nitrogen acting against Teflon expulsion bladders. The propellants are hypergolic and no ignition system is required [e.g., Boeing Company, 1968a]. The engine is mounted on two-axis gimbals with electrical-mechanical actuators providing thrust directional control during engine operations. A central nitrogen storage tank provides the gas required to expel the propellants in the velocity control system and the gas for the attitude control thrusters. The nominal gas pressure for Mission V was serviced with 16.80 pounds of nitrogen at 4,100 psi. The specified propellant load provides a nominal velocity change capability of 1,017 meters per second at an oxidizer-to-fuel ratio of 2.0 and a propellant expulsion efficiency of 98%. On Mission V the four engine burn periods imported a total of 922.41 meters per second of the calculated 1,000.9 meters per second capability and the remaining capability was retained for use during the extended mission. Activation of the propellant heaters was not required during the mission because the system temperatures were generally in the range of 50 to 75oF. 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. The subsystem provided both three-axis stabilization and attitude control [e.g., Boeing Company, 1968a]. 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, although the Canopus tracker operation was significantly restricted by the continuous solar illumination and was only operated in the open-loop mode. 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. Multi-axis spacecraft maneuvers were required to support the four velocity change maneuvers and to orient the spacecraft for the 75 photo sequences. Three-axis maneuvers were executed to support velocity changes and perilune photography while two-axis maneuvers were used for apolune photos. Attitude control was essential for this extremely complex photo mission. Spacecraft maneuvers for nearside photography were based on the requirements that the camera axis point directly at the specified target position at the midpoint of a sequence of exposures and that the image motion compensation be provided by aligning the spacecraft X-axis in the direction of travel. All commands received from the command decoder were properly acted on by the flight programmer, including real-time and stored program commands. The Attitude Control Subsystem supported all operational functions using 472 single-axis maneuvers during its 27-day photo mission. During two-thirds of the mission, the spacecraft was pitched 30 to 54 degrees away from the sunline for about 66% of the mission to maintain temperature control due to the continuous solar illumination. All but two of the 91 Sun acquisitions were accomplished In several instances, reverse photo and thermal pitch-off maneuvers were combined into one maneuver to conserve nitrogen gas. This mission used control procedures developed during the previous missions. Sun sensor operation was also satisfactory. 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, 1968a, b]. 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. A booster regulator was incorporated into the Lunar Orbiter V power subsystem to ensure proper voltages to the photo subsystem during all photo maneuvers, including off-Sun operation (all previous mission required on-Sun operation) 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 V was completely satisfactory. The booster regulator installed for this mission maintained the nominal voltages to the photo subsystem even when the solar panels were oriented at greater than 90 degrees to the sunline. Battery power was used prior to solar panel deployment; during the mid-course, injection, and orbit transfer maneuvers; and during some photo maneuvers. Otherwise electrical power came directly from the solar panels. Thermal Control Subsystem ------------------------- The thermal environment of Mission V was more severe than previous missions and required additional heat rejection capability [e.g., Boeing Company, 1968a, b]. The objective of spacecraft thermal control subsystem was to maintain average spacecraft temperature in the thermal barrier within the range of 35 to 85 degrees F. The thermal control subsystem of the Lunar Orbiter spacecraft is a passive system mounted on a Sun-oriented equipment mounting deck (EMD). Thermal control was maintained by a multilayer aluminized Mylar and Dacron thermal blanket which enshrouded the main bus, special paint, insulation, and small heaters. Eighty percent of the equipment mounting deck EMD was coated with a low-solar-absorbance paint. Twenty percent of the EMD was covered with a geometrical pattern of mirrors to reflect solar energy and reduce the thermal heat dissipation problems. An additional 72 one-inch square mirrors were added under the photo subsystem and adjacent to the TWTA where additional heat rejection was required. This in addition to the 70 mirrors added in a geometric pattern on the EMD for Mission IV. 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. Like Mission IV, the spacecraft was constantly illuminated by the Sun and required a partial off-Sun line orientation for temperature control. To reduce the degradation to the spacecraft protective paint on the outer face of the equipment mounting deck due to continual solar illumination, the spacecraft was flown off the sunline for long periods. Lunar Orbiter V encountered more heat from the continual solar illumination and the additional radiated and reflected heat from the lunar surface and the low perilune. Adding mirrors to the equipment mounting deck and pitching the spacecraft off the sunline satisfactorily controlled the spacecraft temperature. Photographic Subsystem ---------------------- Photography was at least an order of magnitude improved in resolution over Mission IV in numerous nearside locations [e.g., Beeler and Michlovitz, 1969; Boeing Company, 1968a, b; Hansen, 1970; Kosofsky and El-Baz, 1970; Anderson and Miller, 1971]. Farside photography provided coverage of essentially all areas not covered by the previous lunar orbiter missions. This experiment consisted of a dual-lens camera system designed to satisfy the primary mission objective of providing photographic information for the evaluation of Apollo and Surveyor landing sites. An 80-mm lens system was used to obtain Medium-Resolution (MR) photos, and a 610-mm lens system was used for High-Resolution (HR) photos. The two separate lens, shutter, and platen systems utilized the same film supply and recorded imagery simultaneously in adjacent areas on 70-mm film. Automatic sequences of 1, 4, 8, or 16 photos could be obtained. At an altitude of 96 km, which was approximately the perilune height, the HR system photographed a 10- to 35-km area of the lunar surface which was centered on a 70- by 90-km area photographed by the MR system. Resolutions were 2 and 20 m, respectively. At apolune, on the moon's farside at about 6000-km altitude, the areas photographed were correspondingly larger. The film was Bimat processed on board and optically scanned, and the resulting video signal was telemetered to ground stations. Film density readout was accomplished by a high-intensity light beam focused to a 6.5-micron-diameter spot on the spacecraft film. The spot scanner swept 2.67 mm in the long dimension of the spacecraft film. This process was repeated 286 times for each millimeter of film scanned. The raster was composed of 2.67- by 65-mm scan lines along the film. The video signal received at the ground station was recorded on magnetic tape and also fed to Ground Reconstruction Equipment (GRE), which reproduced the portion of the image contained in one raster on a 35-mm film positive framelet. Over 26 framelets were required for a complete MR photograph and 86 for a complete HR image. Of the 213 simultaneous exposures obtained, all were read out satisfactorily. Experiment performance was nominal until the final readout. [Abstracted from: http://nssdc.gsfc.nasa.gov/nmc/masterCatalog.do?sc=1967- 075A&ex=01] Photo subsystem performance was satisfactory during the photographic mission. All photos taken were processed and read out. The photo subsystem of Lunar Orbiter V produced the best quality photographs of the five missions and only one of the 426 photographs taken was not recovered. 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; Boeing Company, 1968a, b]. 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 communications subsystem was employed in the same manner as during Mission IV and performed satisfactorily in all phases. To produce the minimum interference with communication from Lunar Orbiters II and II, the VCO was increased by an average of 330 Hz to offset the ground transmission frequency [e.g., Boeing Company, 1968a, b]. 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 V 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. 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." END_OBJECT = INSTRUMENT_HOST_INFORMATION OBJECT = INSTRUMENT_HOST_REFERENCE_INFO REFERENCE_KEY_ID = "ANDERSON&MILLER1971" END_OBJECT = INSTRUMENT_HOST_REFERENCE_INFO OBJECT = INSTRUMENT_HOST_REFERENCE_INFO REFERENCE_KEY_ID = "BEELER&MICHLOVITZ1969" END_OBJECT = INSTRUMENT_HOST_REFERENCE_INFO OBJECT = INSTRUMENT_HOST_REFERENCE_INFO REFERENCE_KEY_ID = "BOEINGCO1968A" END_OBJECT = INSTRUMENT_HOST_REFERENCE_INFO OBJECT = INSTRUMENT_HOST_REFERENCE_INFO REFERENCE_KEY_ID = "BOEINGCO1968B" END_OBJECT = INSTRUMENT_HOST_REFERENCE_INFO OBJECT = INSTRUMENT_HOST_REFERENCE_INFO REFERENCE_KEY_ID = "BOWKER&HUGHES1971" END_OBJECT = INSTRUMENT_HOST_REFERENCE_INFO OBJECT = INSTRUMENT_HOST_REFERENCE_INFO REFERENCE_KEY_ID = "BUNDICKETAL1965" END_OBJECT = INSTRUMENT_HOST_REFERENCE_INFO OBJECT = INSTRUMENT_HOST_REFERENCE_INFO REFERENCE_KEY_ID = "BYERS1977" END_OBJECT = INSTRUMENT_HOST_REFERENCE_INFO OBJECT = INSTRUMENT_HOST_REFERENCE_INFO REFERENCE_KEY_ID = "HANSEN1970" END_OBJECT = INSTRUMENT_HOST_REFERENCE_INFO OBJECT = INSTRUMENT_HOST_REFERENCE_INFO REFERENCE_KEY_ID = "KOSOFSKY&ELBAZ1970" END_OBJECT = INSTRUMENT_HOST_REFERENCE_INFO END_OBJECT = INSTRUMENT_HOST END References Cited (remove these prior to completion of document. Note: The Boeing refs below are specific to LO-III and are not the same as 'Boeing Company, 1968a, b' in other docs. These may need to be changed to conform to what's in the 'volinfo.txt'. Anderson, A.T. and E.R. Miller, 1971, Lunar Orbiter Photographic Support Data, NSSDC-71-13, 572 pp. Beeler, M. and K. Michlovitz, 1969, Lunar Orbiter Photographic Data, Data Users' Note, NSSDC 69-05, NASA Goddard Space Flight Center, 44 pp. Boeing Company, The, 1968a (e?), Lunar Orbiter V Photography, NASA CR-1094, 244 pp. Boeing Company, The, 1968b (f?), Lunar Orbiter V Photographic Mission Summary, NASA CR-1095, 161 pp. Bowker, D.E. and J.K. Hughes, 1971, Lunar Orbiter Photographic Atlas of the Moon, NASA SP-206. Bundick, W.T., C.H. Green, and E.A. Brummer, 1965, The Lunar Orbiter Telecommunications System, NASA-TM-X-56768, 23 pp. Byers, B.K., 1977, Destination Moon: A history of the Lunar Orbiter program, NASA Headquarters, TM X-3487, Wash, D.C. Hansen, Thomas P., 1970, Guide to Lunar Orbiter Photographs, NASA SP-242. Kosofsky, Leon J. and Farouk El-Baz, 1970, The Moon as Viewed by Lunar Orbiter, NASA SP-200.