The Syncom III Mission and Spacecraft

 Note: This material is taken from NASA Technical Report TR R-252, Syncom Engineering Report Volume II with some editing.

This Syncom Engineering Report, is based on material furnished by the Hughes Aircraft Company and the U. S. Army Satellite Communications Agency, and will cover the launch of the Syncom III satellite, its performance during the first 100 days in orbit, televising of the 1964 Summer Olympic Games by means of the satellite, and various communication tests conducted with it. Syncom III is one of three communications satellites designed and built by Hughes for the Goddard Space Flight Center which have been launched into synchronous orbit. Syncom I was successfully launched in February 1963, but radio contact with the spacecraft was lost shortly after the apogee motor was fired, probably because of an explosion of a nitrogen control system tank. Syncom II was successfully launched in July 1963, becoming the world’s first operational synchronous satellite. This satellite was eventually placed at an area of low perturbation forces over the Indian Ocean after all control system propellant had been expended. From this position it has provided communication links between the Far East, Africa and Europe.

After the launch of Syncom II, the following modifications were made in the Syncom spacecraft:

  1. The nitrogen control unit was replaced with a second hydrogen peroxide control unit.
  2. The apogee motor timer was deleted and redundant provisions were made for a firing by ground command.
  3. Four temperature sensors were provided instead of the previous two sensors.
  4. The standby battery was eliminated.
  5. The P-N type solar cells were replaced by N-P cells and the 0.006-inch cover glass was replaced by 0.012-inch fused quartz covering.
  6. The 500-kc bandpass communications channel was eliminated and replaced by a 10 Mc bandwidth channel for television tests with a 50-kc option for small station testing.

Prior to the Syncom III launch, booster thrust limitations precluded any attempt to reduce the inclination of the Syncom orbit with maneuvers during the boost phase of launch. As a result, Syncom II, although in a synchronous orbit, moves 32″ north and south of the Equator daily. The ultimate objective of synchronous communications advocates has been to place a satellite into synchronous orbit in the equatorial plane. The satellite would then appear to remain stationary and would permit the use of fixed ground antennas without the expense of costly tracking systems. The achievement of this objective by Syncom III became possible with the development in early 1964 of the higher powered Thrust Augmented Delta launch vehicle that could provide enough thrust to permit in flight maneuvers to decrease the orbit inclination.

With the Summer Olympic Games of 1964 scheduled to be held in Japan in early October, the use of Syncom III to present live television coverage of the Olympic Games for the American public became a second launch object.

Syncom III was launched on 19 August 1964 from Pad 17A at Cape Kennedy, Florida. The launch was near-perfect. The maneuver to reorient the third stage for firing at the Equator crossing was also near-perfect. Third stage burn was good, but coning was experienced after burnout. Syncom III separated with a 14-degree attitude error. Seventeen hours and fifteen minutes later, as the spacecraft approached its second apogee over South America, this error was corrected and the spacecraft was oriented to the proper attitude in preparation for apogee motor firing. At third apogee, 29 hours and 2 minutes after liftoff, at a point above the Equator in Borneo, the apogee motor was fired. Syncom III went into synchronous equatorial orbit and later was maneuvered to a position above the intersection of the Equator and the International Date Line.

Since the launch of Syncom III, the following achievements have been recorded, demonstrating the operational and economic advantages of the synchronous communication satellite:

  1. The satellite was put into synchronous equatorial (stationary) orbit over the International Date Line. Ground stations are able to acquire the satellite and lock their antennas in place.
  2. The first 24 hour-per-day, 7 day-per-week reliable communications network has been established across the Pacific Ocean.
  3. Live television of the Olympic Games from Japan was a technical success.
  4. The first communications through an orbiting satellite to a commercial airliner in flight was demonstrated.
  5. The narrow bandpass transponder (50 kc) has provided small-station communications capabilities.

The performance of Syncom III has been excellent and no malfunctions have occurred. The satellite has been in almost constant operation since lift off. The only times the transponders have been off was during apogee motor firing and for a few hours each day when the satellite was in the eclipse season.

 SPACECRAFT DESCRIPTION

The Syncom spacecraft shown in Figures II-1 and II-21 is a spin-stabilized vehicle incorporating electronic, propulsion, and control elements, plus an electrical power supply and a structure. These are described briefly in the following paragraphs.Syncom S:CCommunication Subsystem

The communication subsystem is a redundant, frequency-translation, active-repeater system. Incoming signals from either one or two ground stations at a frequency of approximately 7400 Mc are received by an antenna with a pattern which is symmetrical about the satellite spin axis.

These signals are supplied to two receivers, only one of which is operating at any one time, the desired receiver being selected by command.

Both receivers are wide band, one with a bandwidth of 4.5 Mc between half-power points and the other with a bandwidth of 13.5 Mc between half-power points. In addition, the bandwidth of the 13-Mc receiver can be switched to 50 kc.  Each receiver consists of a mixer, a local oscillator, an IF amplifier, and a limiter amplifier.

When simultaneous two-way, narrow-band communication takes place (duplex operation), the two signal channels are passed through the wide-band amplifier and its limiter. They then modulate the transmitter and are transmitted with power levels approximately proportional to their received signal level.

 Command Subsystem

The command subsystem consists of receivers, decoders, and an antenna unit shared with the telemetry subsystem. The antennas consist of two pairs of whips connected through baluns to the two inputs of a hybrid. The two outputs of the hybrid correspond to the two polarization modes of the whips acting as a turnstile system. Each output is connected to a diplexer. Each diplexer provides 148-Mc command signals to a receiver and accepts 136-Mc telemetry transmission from a telemetry transmitter.

The two command receivers are identical, parallel units each with mixer, IF amplifier, and AM detector. The detector outputs of the two receivers provide audio output tones recovered from the modulation on the command transmission from the ground. Each command receiver is associated with one of the two redundant command decoders. Either receiver/decoder can exercise complete command control of the spacecraft.

Telemetry Subsystem

The telemetry subsystem consists of the antenna (shared with the command subsystem), two transmitters, each with its associated encoder, and the signal conversion elements. The 1.25- watt, 136-Mc transmitter is phase-modulated by a subcarrier which, in turn, is frequency- modulated by a time-division-multiplexed modulator that samples the amplitude of the various sensor signals. Certain critical control signals bypass the time-division-multiplexed modulator and are permitted to phase-modulate the telemetry transmitter directly.

Each transmitter and its associated encoder is one of a redundant pair, each pair operating at a unique frequency. Only one of the transmitter-encoder subsystems is permitted to function at one time, the power to the other subsystem being automatically turned off when the one is turned on. On command, encoder 2 may be disconnected, thereby removing the telemetry modulation and leaving the telemetry carrier to serve as a tracking beacon for the Minitracknetwork.

 Orbit Injection Propulsion

The Syncom orbit injection propulsion subsystem supplies the boost necessary to inject the spacecraft into a nominally synchronous, circular orbit after the vehicle has reached the apogee of a transfer orbit at the required altitude. The spacecraft is launched into the transfer orbit by the Thrust-Augmented Delta vehicle.

The propulsion subsystem consists of a single solid-propellant rocket motor. This motor is required to impart a velocity increment of 4696 fps to the spacecraft, which initially weighs 144.77 pounds.         The following parameters apply to this motor:

Specific impulse                       274.2 seconds

Propellant weight                     60.5 pounds

Case and nozzle weight          10.5 pounds

(including provision for attachment)

Motor weight                            71.0 pounds

Diameter                                  12.0 inches

Payload                                    75.8 pounds

The required performance and objectives given above are met by the JPL rocket engine designated the Starfinder.

Control Subsystem

The control subsystem consists of the components necessary to establish the desired longitude, to maintain a synchronous orbital velocity, and to orient the satellite spin axis from boost attitude to operating attitude. The subsystem consists of two pulsed-jet hydrogen peroxide propulsion units for velocity and orientation control, solar sensors, and control circuits.  An accelerometer for indicating firing is part of the control subsystem.

Electrical Power Subsystem

The electrical power subsystem consists of silicon solar cells, a nickel-cadmium battery, combined voltage regulators and switches. The subsystem is capable of supplying approximately 31 watts without drain on the battery when the satellite is not shadowed by the earth. The solar cells are arrayed on the external cylindrical surface. In the operating configuration, the sunline will be within 25 degrees of normal to the axis of the cylinder, a condition met by suitable choice of launch time.

Structure

The spacecraft structure includes a central, circular member with the separation flange for the Delta third stage at one end and attachment fittings for the apogee motor at the other end. A circularly symmetric bulkhead with reinforcing ribs on the separation side is attached to the member. The electronic units, gas tanks, and four solar cell panels are mounted to the bulkhead. The separation end of the central circular member carries the folding communication antenna. The whip antennas are attached at the apogee motor end of the solar cell panels.  Both ends are closed by thermal shields; the shield at the antenna end also serves as an antenna ground plane.Syncom 3V ASyncom 3V BSyncom 3V C

 

 

Comments are closed.