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 = "LO4" OBJECT = INSTRUMENT_HOST_INFORMATION INSTRUMENT_HOST_NAME = "LUNAR ORBITER IV" INSTRUMENT_HOST_TYPE = "SPACECRAFT" INSTRUMENT_HOST_DESC = " LUNAR ORBITER IV 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=1967-041A). 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 IV 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. The Lunar Orbiter IV mission was the first orbital photographic mapping mission of a celestial body other than Earth. The photo data was used to redefine many of the planned sites for Mission V to optimize and maximize the scientific data requirements. All of the objectives of the Lunar Orbiter IV mission were accomplished [e.g., Boeing Company, 1968a]. LUNAR ORBITER IV 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. 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. Battery power was rarely required during this mission because the high orbit kept the spacecraft from experiencing Earth occultation periods. All components comprising the structure, thermal control, wiring, and mechanisms - except for the camera thermal door - operated properly during the mission. 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 provided the velocity change capability for mid-course correction, lunar orbit injection, and orbit adjustment, as required. The spacecraft included a 100-pound thrust, gimbaled, liquid-fuel rocket engine. The propulsion system used a radiation-cooled bipropellant liquid rocket engine with nitrogen tetroxide as the oxidizer and Aerozine-50 as the fuel. The hypergolic propellants were expelled from the tanks by pressurized nitrogen acting against the Teflon expulsion bladders [e.g., Boeing Company, 1968a]. Due to the change in the mission objective after initial programming, a large midcourse maneuver velocity change was required to rotate the lunar injection point from a 21o descending node orbit to an 85o ascending node orbit. The velocity control rocket engine operated for 501.9 seconds, with a velocity reduction of 659.6 meters per second. The velocity control subsystem and operation was excellent during each of the two propulsion maneuvers both in midcourse and orbit injection. Attitude Control Subsystem -------------------------- The Attitude Control Subsystem provided both three-axis stabilization and attitude control [e.g., Boeing Company, 1968a]. These were provided by four one-lb nitrogen gas jets mounted on the periphery of the engine deck. The attitude reference for yaw and pitch was provided by five sun sensors about the spacecraft to provide spherical coverage and ensure Sun acquisition and lock-on to assist with realignment of solar panels. 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. The inertial reference unit maintains the spacecraft attitude with three gyros providing appropriate rate or angular deviation information to maintain proper attitude and position control. A linear accelerometer provided velocity change information to the flight programmer during firing of the velocity control engine. The 85-degree orbit resulted in the spacecraft being illuminated by the Sun for the entire mission, which reduced the use of the Canopus tracker significantly. Thermal pitch-off maneuvers were needed to dissipate heat from the spacecraft. During the mission there were a total of 83 Sun acquisitions, of which 81 were accomplished in the narrow dead-band attitude control mode. Multi-axis spacecraft maneuvers were required to perform the two velocity maneuvers and orient the spacecraft for the photo sequences. The increased complexity and activity associated with the mapping mission required 586 maneuvers performed, during the 27 day mission, rather than the 284 to 383 of previous missions. Real-time operational procedures were able to minimize the effects due to the problems that occurred during the photo mission. Subsystem performance was generally satisfactory in supporting all mission objectives and there was no difficulty in commanding accurate orientation of the spacecraft to support photography. 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]. All electrical power required and used by the spacecraft was generated by the solar cells. The number of solar cells provided allowed for the possibility that some cells might fail or be damaged by micrometeorites during the mission. The batteries were 28-volt, rechargeable, nickel-cadmium storage batteries. The high polar orbit of Lunar Orbiter IV had several effects. With the solar panels were exposed to the Sun during the nearly entire mission, battery power was required only prior to solar panel deployment and during midcourse and injection maneuvers. Also, a modified charge-controller component was needed to reduce the rate of charge in the power system. Power regulators and controllers protected the solar panels, battery, and spacecraft subsystems from unusual power fluctuations. Performance of the subsystem was satisfactory in all respects during the entire mission. Thermal Control Subsystem ------------------------- The objective of the thermal control system was to maintain the average spacecraft temperature within the thermal barrier to be in the range of 35 to 85 degrees F [e.g., Boeing Company, 1968a, b]. This is a passive system was 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. Heat generated by equipment was conducted to the EMD, where it was radiated into the space environment. To help with heat balance, the exterior surface of the equipment mounting deck was painted with zinc-oxide pigment, silicone based paint with high emissivity in the infrared and low absorption in the wavelengths that contain most of the Sun's emitted heat. Mirrors arranged in a geometric pattern reflected solar energy along 20% of the equipment mounting deck surface. 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 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. Farside photographs, from Lunar Orbiter IV as well as the other LO missions, depended on commands from the Flight Programmer. With a command storage capacity of 16 hours, eight hours more than the amount of time the spacecraft would be out of communications with any of the DSN ground receiving stations, the spacecraft could take photos of the farside without further command from Earth. The mission provided the first detailed near-vertical photographs of the moon and improved visibility of polar regions and limb areas, successfully altered cislunar trajectory by the midcourse maneuver, provided data from which to determine the lunar mathematical model coefficients for an 85o orbit inclination, provided photo data for Lunar Orbiter V mission. Unlike the other missions, the high-altitude orbit negated the need for image-motion compensation. In all only 30 of 163 exposed frames were not recovered. Photo subsystem performance was satisfactory for two-thirds of the active photography portion of the mission [e.g., Boeing Company, 1968a, b]. 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. Lunar Orbiter IV was one of three spacecraft orbiting the Moon and operating on the same frequency. To reliably track and communicate with LO IV and avoid commanding the other spacecraft, an offset track synchronization frequency was employed based on the best lock frequency of the transponder. The communications subsystem was operated differently during Mission IV: The traveling-wave-tube amplifier was turned on for 90% of the time in lunar orbit and was left on to support Manned Space Flight Network tests during the extended mission. The ground transmission frequency was offset (about 330 Hz) to produce minimum interference with Lunar Orbiters II and III. Overall performance of the communications subsystem in high-power, low-power commends, and ranging and/or tracking modes was satisfactory and the performance of the subsystem was satisfactory in all phases [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 IV 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 Space Flight Operations Facility (SFOF) provided the mission control center and the facilities to process and display data to support operational mission control. The telemetry system samples the output of sensors within the various spacecraft subsystems. Normal telemetry data channels including information as temperatures, pressures, voltages, currents, and error signals. Special instrumentation included micrometeoroid detectors, scintillation counter dosimeters, and the associated logic. The traveling wave tube amplifier transmitted wide-band video data and telemetry during readout. Initial receipt of telemetry data via the DSN primary sites occurred 32 minutes after liftoff at the Deep Space Station at Johannesburg, Africa. The initial track was lost and the Woomera Deep Space Station acquired the spacecraft 45 minutes after launch." 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