The Surveyor mission as defined by NASA and JPL in 1960 was to soft land a package of instruments on the moon so that this payload would operate after landing. In January 1961 when Hughes was awarded the contract for Surveyor no spacecraft had safely landed on the moon. The first attempted lunar missions took place in 1958; four Pioneer missions by the United States attempted to orbit the moon and three Soviet Luna missions attempted to impact the moon—none were successful. The Soviet Luna 2 was the first spacecraft to impact the moon in September 1959 followed by Luna 3 in October that returned images of the far side of the moon. It was against this background that JPL and Hughes Aircraft set out to design a spacecraft that could achieve this goal.
This Surveyor mission description will be documented in four separately posted reports: 1. Surveyor Mission Description, 2. Free Flight Trajectory Design, 3. Midcourse Guidance and 4. Terminal Guidance and Descent.
1. Surveyor Mission Description
All seven Surveyor missions were successfully launched by an Atlas-Centaur utilizing Complex 36 of the Kennedy Space Center. Two different launch modes were used—direct ascent using a continuous Centaur burn and a parking orbit ascent during which the Centaur shuts down and then restarts after a coast period. When NASA’s Lewis Research Center (now Glenn) took over the Centaur development program it was decided to delay development of the parking orbit capability until the other Centaur issues had been resolved. The initial Centaur design wasn’t able to provide the required propellant settling during the coast period to allow an engine restart. For the historical record three launches used the direct ascent mode: Surveyor I (5/30/66), Surveyor II (9/20/66), and Surveyor IV (7/14/67); and four used the parking orbit mode: Surveyor III (4/17/67), Surveyor V (9/8/67), Surveyor VI (11/7/67) and Surveyor VII (1/7/68).
Note: The following edited mission summary was originally published in Hughes SSD 56028R Surveyor Trajectory Characteristics, December 1965.
The Atlas-Centaur boosts the spacecraft to into a lunar transfer trajectory that will reach the moon in approximately 66 hours. Following separation from the spacecraft the Centaur is turned 180 degrees and the remaining propellant is expended so that the vehicle will not impact the moon or interfere with Surveyor operations. The Surveyor after separation deploys the landing legs, omni antenna, and solar panel to its cruise position. Following these deployments the attitude jets stabilize any separation tip-off rates and a sun acquisition sequence is initiated.
Approximately 6 hours after launch, a command is given to acquire the star Canopus, by means of a roll maneuver, while maintaining sun lock. After Canopus acquisition is signaled by the spacecraft, a verification procedure is performed to ensure the sensor has locked on Canopus and not another star or planet. At this point Surveyor is fully attitude stabilized.
During the initial coast phase engineering telemetry is gathered to assess the health of the spacecraft and tracking data is utilized to accurately determine its orbit. Nominally at 15 to 20 hours after launch, with visibility from the Goldstone DSIF, a midcourse maneuver is performed using the three throttleable vernier engines. This maneuver corrects Atlas-Centaur injection errors, adjusts the location of the landing site if desired, and assures that the initial conditions for terminal descent provide high probability for soft landing. On command, the spacecraft performs a series of attitude changes that aligns the thrust axis in the required direction. The three vernier engines are then ignited and burn at a fixed acceleration level for a commanded time to provide the desired velocity increment.
Following the completion of the midcourse maneuver, the spacecraft is returned to the cruise attitude locked on the sun and Canopus. Engineering telemetry and tracking data are gathered by the DSIF and furnished to the SFOF for analysis of spacecraft status. Further orbit determination allows evaluation of the midcourse maneuver and computation of final terminal maneuver parameters.
As the spacecraft, in view of the Goldstone DSIF, approaches the moon 66 hours after launch, the thrust axis is aligned with the velocity vector. The high-gain antenna is pointed at the earth if approach television pictures of the lunar surface are desired. At a slant range of 60 miles, the altitude marking radar generates a signal that, after a suitable time delay established by ground command, initiates the ignition of the vernier engines that are used to control the spacecraft attitude during the main retro burning. One second after vernier ignition, the solid main retro motor is ignited. After 41 seconds thrust decay is sensed by an acceleration switch that initiates a time delay of 12 seconds and the empty main retro case is jettisoned.
Shortly after separation, vernier engine thrust is reduced to 0.9 lunar g’s. The radar altimeter and doppler velocity sensor (RADVS) senses the lunar surface and provides measurements of range and velocity to the flight control system, and the thrust axis is aligned with the velocity vector. When the RADVS senses that the pre-programmed range-velocity contour has been reached, the thrust is increased to an acceleration level such that the spacecraft descends along this contour until a speed of 10 fps is reached. The dynamic character of the descent trajectory (a “gravity turn trajectory”) is such that when the spacecraft reaches the 10 fps point, the thrust attitude will be nearly aligned to the vertical. The vehicle attitude is then inertially fixed and the maximum acceleration is held down to an altitude of 13 feet when the engines are shut off and the vehicle free falls to the surface with a landing speed of 13 feet/second.
After landing on the surface of the moon, the spacecraft is commanded to align the solar panel to the sun-line and the high grain antenna to earth. Subsequent commands connect one of the transmitters to the to the planar array antenna and switch that mode (high or low power) that will provide the necessary communications bandwidth consistent with the operation being performed, the power available fro the solar panel and battery, and the thermal restrictions imposed by the compartment dissipating capabilities. At this point, any of various engineering or scientific experiment sequences may commence.