DESTINATION MOON: A History of the Lunar Orbiter Program
 
 
CHAPTER VI: THE LUNAR ORBITER SPACECRAFT
 
A General Description
 
 
 
[111] Before surveying the design and development phases of the Lunar Orbiter Program, it will be useful to describe the spacecraft which Boeing built for Langley. In the final design the Boeing Orbiter weighed about 385 kilograms and was 1.7 meters tall and 1.5 meters in diameter at its base, without including the solar panels and the antennas. Structurally the spacecraft had three decks supported by trusses and an arch. On the largest deck the main equipment was mounted: batteries, transponder, flight programmer, photographic system, inertial reference unit (IRU), Canopus star tracker, command decoder, multiplex encoder., and the traveling-wave-tube amplifier (TWTA), together with smaller units. Four solar panels and two antennas extended from the perimeter of this equipment deck.1
 
Above it, the middle deck supported the velocity control engine (the 100-pound-thrust Marquardt rocket motor), the fuel tanks, the oxidizer tank for the velocity control engine, the coarse Sun sensor, and the micrometeoroid [112] detectors. Above this the third deck contained the heat shield to protect the spacecraft from the heat generated by the firing of the velocity control engine. In addition the four attitude control thrusters were mounted on its perimeter. This uppermost deck was part of the engine module, which could be detached for test purposes. Directly under the engine was the high-pressure nitrogen tank, which provided pressure to feed fuel to the velocity control engine and to operate the attitude control thrusters.2 This tank was one of the critical units; if anything caused it to lose pressure, the spacecraft could not maneuver, and an entire mission could be ruined.
 
These and other items of spacecraft equipment formed subsystems of the whole spacecraft system. Working together they performed the Lunar Orbiter mission. The Eastman Kodak photographic subsystem has previously been described.3 Electrical power was provided by a power system which operated in two modes: 1) solar panels converted solar radiation into electric current, and 2) batteries powered the spacecraft systems for short periods of occultation from the Sun. In periods when the solar panels would receive radiation from the Sun, the power supply would [113] run from the panels through the output voltage regulator to the other spacecraft systems (mode 1). This happened for the major part of the mission. At the same time power generated by the panels would also be directed into the battery charge controller, and from there a charging current would flow into the batteries as they could accept it. When no sunlight fell on the panels, the batteries would supply power to the output voltage regulator, and this would direct its flow to the spacecraft subsystems (mode 2).4 In addition the power system had regulators and controllers to reduce unusual fluctuations to a minimum and enough solar cells to allow micrometeoroid damage to some without dangerous reduction in the capacity of the solar panels to generate electricity.
 
The attitude control subsystem served as the navigator for Lunar Orbiter during an entire mission. Composed of Sun sensors, the Canopus sensor, the inertial reference unit, and the thrusters, the system controlled the spacecraft's attitude in space in reference to the Sun, the star Canopus, and the Moon. The Sun sensors would "see" the Sun, produce signals which activated the attitude control thrusters, and these would align the spacecraft's roll axis with the sun. Once this reference was established the spacecraft could maneuver off the reference and the IRU would remember [114] the original reference. If the need arose to move the spacecraft back to that reference, the IRU would signal the thrusters to correct the attitude. However, the IRU simply remembered reference points; it did not establish them.
 
Attitude control was directed by the flight electronics control assembly (FECA) and the Flight Programmer, which received data from all sensors and then informed ground control monitors, who could update the Programmer for future attitude maneuvers. The FECA and the Flight Programmer controlled the spacecraft's attitude around its X (roll), Y (yaw), and Z (pitch) axes by activating the thrusters. They also governed the orientation of the photographic subsystem's camera lenses in relation to the surface of the Moon. Commands from Earth would make the spacecraft rotate through an angle around each axis according to the task to be executed, and the outputs of the gyros in the IRU would tell the Flight Programmer when the new attitude had been achieved. The Flight Programmer would stabilize and maintain the spacecraft in the new attitude relative to the three reference directions, and the IRU would tell it when there was any deviation from the established attitude.5
 
[115] The Atlas-Agena D launch vehicle placed all five of the Lunar Orbiter spacecraft in parking orbits around Earth. The Agena with the spacecraft would remain in the parking orbit until the time to begin the translunar trajectory maneuver in which the Agena, would fire out of Earth orbit toward the Moon. Once the spacecraft separated from the Agena there remained the task of correcting its initial trajectory and then of deboosting it into lunar orbit. The velocity control subsystem held the responsibility for this task and had to execute any changes in trajectory and speed.
 
The heart of the system was a 100-pound-thrust rocket whose hypergolic, fuel and oxidizer ignited when the Flight Programmer commanded the intake valves to open. A burn to change the spacecraft's velocity would then occur and continue until the valves closed. Duration of any burn would be determined by information from the accelerometers in the IRU compared with prestored data in the Flight Programmer. The rocket engine was gimbaled to provide thrust vector control in order to accommodate center-of-gravity offsets and thrust asymmetries. The IRU accelerometers provided inputs for thrust vector control, the purpose of which was to keep the thrust of the velocity control engine through [116] the spacecraft's center of mass.6
 
A nominal mission would provide for two midcourse maneuvers to bring the Orbiter's trajectory precisely in line with an imaginary point where it would be deboosted into orbit around the Moon. At this predetermined point the velocity control subsystem would fire to slow the spacecraft and allow it to go into an initial orbit around the Moon. Ground personnel would then check out the spacecraft's orbital behavior and its various subsystems before making' any decision to transfer to another orbit. Once they found the spacecraft's subsystems to be operating correctly, they would make a decision to inject it into a photographic orbit.7
 
Receiving and transmitting data to and from the spacecraft was the job of the communications subsystem, many of whose components had been flight-proven in the Ranger and the Mariner programs. This complex assembly could operate in four modes: 1) tracking and ranging, 2) command, 3) low power, and 4) high power. The communications system could send and receive data simultaneously while also transponding velocity and ranging signals for the Deep [117] Space Network's tracking system.
 
The spacecraft's low-gain antenna picked up all incoming signals from the NASA-JPL Deep Space Instrumentation Facility stations. Commands from DSIF were routed to the command decoder and stored. The spacecraft would transmit a command from Earth back to Earth for verification before ground controllers sent an "execute" command. Upon receiving the execute command the communications subsystem would advance stored commands from the decoder to the Flight Programmer to be carried out. Photographic data with performance, environmental, and telemetry data would be transmitted to Earth by the high-power mode.8
 
Photographic data were transmitted in a different way than telemetry data were. The spacecraft had two antennas that operated in the S-band at the frequency of 2295 mega-cycles. Normally, when photographic data were transmitted to the ground receiving stations, the communications subsystems operated in the high-power mode and transmitted via the one-meter-diameter parabolic high-gain antenna. Simultaneous transmission of photographic and telemetry data was carried out as follows:
 
[118] The 50-bit/sec telemetry data train is phase modulated onto a 30-kc subcarrier, which is then combined with the video data that have been transformed to a vestigial sideband signal. That signal is created by amplitude modulating the data on a 310-kc subcarrier by means of a double balanced modulator. This suppresses the carrier and produces two equal sidebands. An appropriate filter is then superimposed on the double sideband spectrum, essentially eliminating the upper sideband.
 
Since the missing subcarrier must be reinserted on the ground for the proper detection of the vestigial sideband signal, provision for deriving such a subcarrier signal is made by transmitting a pilot tone of 38-75 kc. That pilot tone is exactly one-eighth of the original 310-ke subcarrier frequency, and is derived from the same crystal oscillator. Multiplying the received pilot tone by 8 in the ground equipment provides a proper subcarrier for reinsertion.9
 
Lunar Orbiter photographic data were never encoded; instead, data were transmitted as frequency-modulated analog signals. All other data from the spacecraft were encoded and sent on the subcarrier frequency as described above. The temperature control subsystem protected all of the spacecraft's other subsystems from the extreme temperature variations of the deep space environment. Heat from the Sun could warm external parts of the spacecraft to 120°C while areas not exposed to solar radiation would cool down to -160°C. These extremes were beyond the temperature [119] levels which most components could endure. The temperature control system established an environment ranging from + 2°C to +30°C for the operation of all subsystems. A few components were exposed to direct sunlight: the four solar panels, the two antennas, the bottom of the equipment deck. The solar panels were designed to withstand temperature variations of +120°C to -160°C without cracking or buckling from severe expansion and contraction over a long period of time.10
 
Beginning at the uppermost deck a heat shield insulated the spacecraft from the rocket engine's heat while the entire area down to the lower deck was enshrouded in a thin-skinned aluminized mylar and dacron thermal blanket that covered all equipment except the Canopus star tracker's lens, the camera thermal door, and the components mentioned above. The bottom of the equipment deck, which faced the Sun most of the time during all five missions,, was coated with a special paint having a high heat emission-absorption ratio. Small electric heaters were installed on the spacecraft inside the thermal blanket to raise the temperature if it fell below +2°C. The arrangement maintained everything under the thermal blanket at an average temperature.11
 
[120] The photographic subsystem had the most rigid temperature restrictions. Film could withstand heat only up to about 50°C, and moisture In the photographic subsystem would condense below 2°C, fogging the camera's two lenses. Eastman Kodak designed the system to be biased cool and warmed with little electric heaters. The "bathtub" housing the system did not touch the equipment deck but was affixed by four legs. Heat transfer between the "bathtub" and the equipment mounting deck was largely radiative, making heat absorption and dissipation a slower, more even process.12
 
One other component of the temperature control system was added after the original design to protect the photo-subsystem. This was the camera thermal door. Thermal tests showed that without any cover over the camera's lenses, the lenses would be more susceptible to extreme temperature variations and stray light leaks inside. The major purpose of the camera thermal door was to reduce or eliminate the possibility that through heating the lenses could expand and alter the focal length so that distortions would result in the photography. The door would also help to control the internal temperature of the photo-subsystem so that it would not become too cold during periods of occultation and allow moisture condensation on the lenses. The door was added as one of the last components of the [121] spacecraft before final design configurations were fixed. It was not part of the Eastman Kodak camera subsystem, and Boeing took the responsibility of designing, fabricating, and testing it.13
 

 
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