Pioneer Venus Mission Operations Support.—Jack Fisher

Charlie Hall and his Pioneer Project team were very experienced in mission operations. They had their own Mission Operations Center at the Ames Research Center in Sunnyvale and had experience dating back to Pioneers 6-9 in the 1960s and Pioneer 10 and 11 that were launched in 1972-73 and flew the first missions to Jupiter and Saturn. They relied upon the JPL DSN for tracking and communicating with the spacecraft, and JPL orbit determination capabilities, but were otherwise entirely self sufficient. They looked forward to their first missions to Venus with the Pioneer Venus Orbiter and Multiprobe.

ARC realized that successful missions operations would depend upon a detailed knowledge of the spacecraft and its characteristics that would have to be provided by the spacecraft contractor. Accordingly ARC issued Contract No. NAS 2-9366 to Hughes Aircraft in 1976 to provide Operational Characteristics Documents for both the Pioneer Venus Orbiter and Multiprobe. I was given responsibility for this contract. My first task was to find someone to take the lead for this monumental task. Since the project team was extremely busy at this time I had to look elsewhere. I consulted Perry Ackerman and he suggested Ken Filetti. I did not know Ken at this time, but I didn’t have a lot of options, so Ken it was,

Ken set to work, developed a plan, and began. The Orbiter Operational Characteristics document came together between September 1976 and June 1977 and was published on February 28, 1978. It describes the orbiter system and interfaces, each subsystem, its operating modes, and mission operations in a total of 963 pages. The authors of this document were Bill Butterworth, Dick Daniel, Bob Drean, Ken Filetti, Jack Fisher, Larry Nowak, Joe Porzucki, Jerry Salvatore, and George Tadler of systems engineering and Fred Barker from propulsion.

The Multiprobe document was prepared between September 1976 and May 1978, and was published on May 22, 1978 in three volumes consisting of 1002 pages and covering the Multiprobe, Large Probe and Small Probes. The authors were again Bill Butterworth, Dick Daniel, Bob Drean, Ken Filetti, Jack Fisher, Larry Nowak, Joe Porzucki, Jerry Salvatore, George Tadler and Fred Barker with the addition of Fred Richardson and Jan Sikola.

Both missions were very successful: the Multiprobe mission was shorter and ended with the successful entries of the Bus, Large Probe and three Small Probes on December 9, 1978 while the Orbiter mission, through careful propellant management, was extended to October 8, 1992, a total of more than 14 years. I am grateful to Ken Filetti and the authors of these documents for their contributions to the success of the Pioneer Venus missions.



Configuration Management—Jack Fisher

Configuration management (CM) is a mystery to almost everyone as it was to me when we starting working on the Pioneer Venus contract in 1974. This was a serious concern as I had been chosen to head systems engineering for the program. My first stop in the quest for understanding was the company management directives. I spent the better part of a week reading and trying to understand these with little success—they were incomprehensible. My next step was to talk to Mal Meredith, my systems engineering mentor—Mal at that time was heading systems engineering for the NASA OSO program and he surely would be able to help me. His answer—don’t worry about it—it’s really very easy to understand and when it becomes important you will pick it up quickly. Needless to say this advice was very welcome.

Configuration management had its origins in the 1950s as the Air Force was developing ballistic missiles. A missile failure would require fixes resulting in a successful launch, but if the fixes were not properly documented they perhaps could not be replicated. CM begins with the formal release of program documents and drawings and requires that modifications be classified and formally controlled so that the product configuration is always known and under control. When the system is to be delivered to the customer its performance and configuration have to be reported.

We started the program and worked the major design issues and advanced into program engineering with the release of specifications and drawings. And this occasioned the need for control of the configuration or configuration management and the appearance of Mike Chekel. The NASA Division had a CM group that Mike headed. We started having weekly Configuration Control Board (CCB) meetings that I was expected to chair. Responsible Engineering Authorities (REAs) would propose changes to modify their subsystem design and the CCB would have to evaluate and cost them making sure that if other subsystems were impacted their REA would evaluate the change. With Mike’s guidance and meticulous record keeping we handled this phase of the program without any major difficulties.

The next hurdle was the Orbiter Pre-Ship Review that was held over a four-day period, February 21-24, 1978. It was chaired by Charlie Hall, the ARC program manager, and attended by representatives of various NASA centers, GDC, the launch vehicle contractor, and the experimenters that provided instruments for the mission. The purpose was to show that the spacecraft met all specification requirements and that we knew precisely what items comprised the spacecraft to be delivered. With Mike Chekel’s support we passed those hurdles with flying colors. Systems engineering prepared a System Performance Assessment Document (SPAD) that summarized spacecraft performance versus the requirements.

We later received a commendation from Charlie Hall, the ARC program manager that thanked us for “a job well done.” The orbiter was shipped to the KSC launch site on March 15 and was launched on its way to Venus on May 20. I left for the Cape in early March and headed up the systems engineering activities at the launch site. I returned briefly to El Segundo for the Multiprobe pre-ship review April 24-29. The Multiprobe was shipped to the launch site on June 5 and was launched on August 8. With Mike Chekel’s guidance I didn’t find that configuration management was all that difficult.

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.


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.


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




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

The coincidence of the availability over the Pacific Hemisphere of Syncom III, with its ability to relay television signals, and the Olympic Games in Japan presented an attractive opportunity to bring the attention of the world to the revolution in communication being wrought by synchronous satellites. The televising of the Olympics live from Japan to the United States had been under consideration both in Japan and the United States since NASA’s decision to locate Syncom III over the Pacific. Much informal investigation and discussion had been carried on among interested agencies.

Near mid-1964, the Space Council reviewed a proposed program for live TV coverage of the Olympics and found it to be in the national interest. The Council suggested the Communications Satellite Corporation (Comsat) as the appropriate U. S. agency to coordinate the program and endorsed the utilization of various U. S. government facilities for the program.

Comsat set about organizing the program as a non-profit endeavor and obtained funding from interested sponsors. Among those expressing interest were the Canadian Broadcasting Corporation and the European Broadcasting Union. The latter company, which distributes program material to the whole of Europe (Eurovision),found that it could bring the events to its audience at a much earlier time (the same day as their occurrence and at prime viewing time)by combining transmission via Syncom III and microwave link from Tokyo to Montreal, Canada, with fast jet aircraft transport from Montreal to Europe. These organizations arranged with Comsat to purchase transmission time for Olympic program material.

The entire program arranged by Comsat is outlined in Comsat’s filing with the Federal Communications Commission (FCC). The report that follows describes the transmissions that took place, the transmitting and receiving equipment used in the program, and lastly, provides a few photographs illustrative of the program material transmitted and a tabulation of receiving, antenna pointing angles attesting to the stationarity of Syncom III’s orbit.


Between 17 September and 24 October, a daily period of three to four hours was devoted to the transmission of video material from the Radio Research Laboratory at Kashima, Japan, via Syncom III to Point Mugu, California. The total number of separate scheduled transmission attempts was approximately 48.  All were successful except the one attempted on 25 September, which was terminated by a typhoon in Japan (that forced stowing of the antennas to prevent damage to them). From 17 September through 9 October, the transmissions consisted primarily of four-hour periods of test patterns and still picture material for station adjustments.

During the period 10 October through 24 October, approx-imately three hours of Olympic Games video program material (from video tape) was transmitted daily from Kashima to Point Mugu. (Voice was transmitted via cable and voice and sound were synchronized by addition of time delay to the voice circuits.) Some material, including the opening day ceremonies in Tokyo, were shown (via microwave relay link) in real time in the eastern United States, Figure VI-1 shows the routing of signals in their passage about the world.Olympic Signal Routing        The quality of the video received on all transmissions to Point Mugu was outstanding as judged by technical and subjective measurement standards.  On most viewers’ sets it was not distinguishable in quality from most other program material, though it was reprocessed many times after its initial receipt at Point Mugu with the normal degradations attendant on such processing steps.

The schedule required by the European Broadcasting Union to provide effective exploitation of Syncom-relayed material was governed by the relative geographic position of Europe and Japan and by the time required for editing the material and transporting it to Europe. These factors, along with the restraints imposed by other users of the satellite, dictated a time of transmission of the material through the satellite which coincided with the eclipse time of Syncom III. It was necessary to interrupt transmissions during the eclipse period to allow the satellite to be turned off. No deteriorating effects of the shadow transit on the relayed video were observed.

The time of day did not affect the quality of the video or the noise levels of the receiving system. The test periods were used for acquisition of the satellite (primarily at the transmitting end where pointing tests were conducted with both the receiving and transmit- ting antennas to obtain optimum illumination of Syncom by the transmitter) and for adjusting the modulation parameters. Signal-to-noise         and carrier-to-noise measurements were made during each test period.

The schedule was adhered to 100 percent of the time, except for five minutes of program material and ten minutes of test time lost on 14 October due to a failure in the transmitter. All transmitted video material was judged acceptable by the exhibiting agencies. This performance attests to the dependability of communication by synchronous-orbit satellites.

Section 10-Intelsat—Compendium of Satellite Communications Programs NASA X-751-73-178 June 1973

10.1 Program Description

The International Telecommunications Satellite Consortium (Intelsat) is a partnership initially established between 14 member nations in 1964 for the purpose of providing global commercial tele-communications via satellite.         Since that time, the organization has expanded to include 79 member nations (as of May 1971), and new applications for membership continue to be received.

From the time it was established in 1964 to the present, Intelsat has produced four generations of satellite and ground systems. Development of the initial satellite, nicknamed Early Bird and later designated Intelsat I, was initiated by Comsat in November 1963. Just a year earlier, in August 1962, the U. S. Congress had passed the Communications Satellite Act, which authorized the creation of a private corporation (Comsat) to instigate the development of a global commercial communications satellite system. At the time when Early Bird’s development began, the Syncom II satellite had just completed demonstrating that reliable communications could be provided through lightweight synchronous satellites. As a result, the Syncom satellite design formed the basis for the Early Bird spacecraft.  Early Bird was launched in early 1965, as indicated in Table 10-1, and by April 22, 1965, had successfully achieved synchronization into the desired geostationary orbit with the satellite located over the Atlantic Ocean. After a period of satellite performance testing, system parameter evaluation using the operating ground stations, terrestrial and satellite circuit lineup, and public demonstrations, commercial operation was initiated on June 28, 1965. The satellite successfully provided commercial communications service between the United States and Europe until it was retired in early 1969. It was reactivated for a brief period later in 1969 when temporary difficulties were encountered with the antenna system of a third generation Intelsat satellite.Early Bird

The second generation of Intelsat spacecraft was designated Intelsat II. Even when the Intelsat I system was under development, it was realized that many of the inherent advantages of space communications could not be exploited. Specifically, its antenna characteristics were such as to embrace only the north-eastern part of North America and the western part of Europe, and it did not allow for simultaneous intercommunication among numerous earth stations. By late 1965, it was recognized that the constraints imposed by these deficiencies would be incompatible with a NASA requirement for multichannel communications in late 1966 among its tracking stations at Carnarvon, Australia, Ascension Island, Canary Island, tracking ships in the Atlantic, Pacific, and Indian Oceans, and the Manned Space Flight Center in Houston, Texas. In the past, these circuits had been carried by HF radio, but for manned space flights, improved communications were desired. Consequently, in the fall of 1965 the development for Intelsat II had begun with the primary goal of satisfying the NASA requirements; excess capacity was to be used for other commercial traffic.  Because of the urgency to satisfy the NASA requirement, the Intelsat II design evolved directly from that of Intelsat I.

Four Intelsat II satellites were produced and launched as indicated in Table 10-2. The first launch occurred in October 1966; but when the satellite’s apogee motor malfunctioned, the spacecraft was left in a highly elliptical inclined orbit, making it unusable for full-time commercial operations.         Subsequent launches in January and September 1967 successfully placed two satellites into operational service over the Pacific Ocean.  A March 1967 launch successfully supplemented the Intelsat II satellite in operation over the Atlantic Ocean. With these three satellites in place, commercial service was available over both the Atlantic and Pacific Oceans.  The Intelsat II satellites continued to meet international commercial communications requirements successfully until third generation replacement satellites allowed them to be retired to the active reserve.Intelsat II           The development of Intelsat III was initiated in 1964 with a design study and followed 2 years later with the award of a contract for the design, development, and fabrication of the necessary spacecraft. Eight Intelsat III satellites were launched between September 1968 and July 1970 as indicated in Table 10-3. Launch failures in September 1968, July 1969, and July 1970 made three of these satellites unusable. A successful launch in December 1968 placed a spacecraft in service over the Atlantic. Some difficulties were encountered when this satellite’s mechanically despun antenna started sticking in mid-1969.  This occurrence made necessary the aforementioned reactivation of Early Bird. The antenna problem was resolved by August 1, 1969, and commercial operations were resumed until a subsequent January 1970 Intelsat Il launch allowed the satellite to be placed in the retired reserve.  In April 1970 another successful Intelsat III launch supplemented the operational capability available over the Atlantic.IntelsatIII        An Intelsat III satellite was first placed into operation over the Pacific in February 1969. This satellite was supplemented by a second spacecraft in May 1969. When the first Pacific Intelsat III lost 6 dB of transponder gain due to an RF receive amplifier malfunction, it was relocated over the Indian Ocean where the traffic requirements were lighter. As a result, four Intelsat III satellites were, as of May 1971, providing global commercial service over the Atlantic, Pacific, and Indian Oceans.

Development of the fourth generation of Intelsat spacecraft, Intelsat IV, began in the latter portion of the 1960s.  These satellites, manufactured by Hughes Aircraft Corporation, have been designed to provide a substantially greater capability to meet the increased global communication needs of the 1970s. The first, in an expected series of eight satellites, was successfully launched by an Atlas Centaur rocket into a geostationary orbit on January 25, 1971.         It was positioned over the Atlantic at 24 5°W longitude. Subsequent Intelsat IVs will be launched for service over the Atlantic, Pacific, and Indian Oceans with an additional two satellites as spares in orbit. Each satellite will have a design life expectancy of about 7 years.

Because the Intelsat program has been a commercial venture, the number of major innovations in equipment and techniques employed has been limited. The intent has been to minimize the risk of spacecraft failure and the exceptional reliability record amassed by the system testifies to the success of this policy. Nevertheless, the Intelsat program has made significant contributions to satellite communications.

Numerous subjective tests with Intelsat I demonstrated conclusively that the round trip time delay and echo due to two-wire user terminations were not insurmountable obstacles to the utilization of synchronous satellites for commercial communications. This was in confirmation of preliminary indications obtained on Project SYNCOM.

The Intelsat II spacecraft demonstrated that tunnel diode amplifiers of operational reliability were available for use as RF receive preamplifiers.         Utilizing these relatively low-noise, high-gain preamplifiers allowed direct RF to RF conversion in a single stage to be employed.  Sufficient spacecraft power and high performance earth terminals allowed these transponders to be designed for linear input/output power transfer characteristics. Additionally, Intelsat II and an expanded ground complex demonstrated the feasibility of extensive multiple accessing of a single satellite transponder by a group of operational ground terminals.

When the wideband Intelsat III satellites were placed into operation, it was necessary to introduce a third generation of earth stations to the system in order to take full advantage of the expanded capabilities of the space subsystem. These terminals employed newly developed 500-MHz bandwidth cooled parametric amplifiers, as well as 500-MHz bandwidth high power traveling-wave tube transmitters, capable of over 6 kW of multi-carrier power.

The more recent Intelsat IV spacecraft have contributed to satellite communications technology by demonstrating fixed narrow beamwidth (i. e., 4.5°) antennas mounted on a mechanically despun platform and a highly channelized satellite repeater (i. e., 12 inde- pendent transponders). The narrow-beam antennas provide coverage to a fixed, relatively restricted area of the earth but the high antenna gain available significantly increases satellite EIRP. The large number of transponders, each having a 36-MHz bandwidth, allows users with substantially different communication requirements to operate independently of each other in separate satellite channels (e. g., television distribution in one channel, high-capacity telephone trunks in another, and low-duty cycle individual voicelinks in still another). Additionally, a fully variable, demand-access, satellite system will be demonstrated for the first time during the period when the Intelsat IV satellites are being placed into operational service. Comsat has developed a single-carrier-per-voice channel PCM-PSK-FDMA demand-access system nicknamed SPADE that will be employed.

Dick Parfitt—Jack Fisher

I don’t remember that in the Pioneer Venus studies prior to the proposal that anyone had a great deal of concern about the spacecraft wire harness. For our proposal effort and program costing exercise the Design Integration organization would have been responsible for the design and construction of the harness and I’m sure that was taken into account in their bid. For my systems engineering bid I was blissfully ignorant of any responsibility for the wire harness.

When we started working under the contract in 1974 I was fortunate enough to recruit several very experienced systems engineers, one of whom was Dick Daniel. Shortly after he started Dick came to tell me that we needed to get Dick Parfitt to do the wire harness Interconnection List (ICL). The ICL is list of connections that the harness must provide and is essentially the starting point for the design of the harness—or the first set of harness requirements and was the responsibility of systems engineering. Well I couldn’t argue with that, but then I learned that Dick Parfitt was in the Commercial Division. I was reluctant until I learned that Dick had done the ICLs for all previous commercial and NASA satellites, with the exception of Surveyor. So Dick Parfitt did the Pioneer Venus ICLs—there were four to be done: the Orbiter, Multiprobe, Large Probe, and the Small Probe. I don’t remember any issues that arose, Dick worked unobtrusively and efficiently and got the job done. I wish I could write more about Dick Parfitt because he was a great asset to SCG. I would invite anyone who has recollections of Dick to add their comments to this post.

ATS Spin Scan Pictures–Bob Farmer

Prior to my joining Hughes in August 1973 and after leaving the US Navy in March of 1966 I was with Bendix Field Engineering Corp (BFEC).  BFEC at that time had the contract with NASA for ground station support of the STADAN, Manned, and Deep Space tracking and operations networks.  I was assigned to the NASA tracking facilities at Goldstone outside of Barstow, CA,  and supported both manned and unmanned space missions.  During my time there my activities were in support of the last four Gemini missions and the complete Apollo program.  In between manned missions we supported various unmanned programs one in particular was the ATS (Applications Technology Satellites) Program.  You mentioned in your writeup the two firsts from ATS-1 and ATS-3 regarding their Spin Scan Cloud Cover camera systems.  I looked through my old (very old) files and found these two pictures.  The black and white is one of the very first from ATS-1 that we processed at Goldstone before sending on to GSFC, in the NE corner you can just make out the outline of Baja California.  The color photo is a repro of the first color picture from ATS-3 that was on an “Atta Boy” we all got for being part of the team.ATS-1 Spin Scan

ATS-3 Spin Scan

Applications Technology Satellite Program

Editor’s Note: This information is taken from a NASA website:—there is no mention of the Hughes people that were involved in this program and I am relying upon our readers to provide comments–pictures also welcome.

The Applications Technology Satellite (ATS) series was conceived of as a follow-on to the successful experimental communications satellites of the early 1960’s with the addition of other technology demonstrations, such as weather observation and investigation of the space environment.

NASA and Hughes had hoped to continue the success of the Syncom project with an Advanced Syncom. They met resistance from some members of Congress who feared that NASA was developing technology for the benefit of a private company, namely Comsat. So the project’s objectives were broadened and it became ATS. The Department of Defense (DoD) influenced NASA to include technology for gravity-gradient stabilization (on ATS-2, ATS-4, and ATS-5) and for medium altitude orbits (ATS-2). All five of the first generation ATS spacecraft carried a C-band transponder with 25 MHz bandwidth capable of relaying 1200 one-way voice circuits or one color television signal.

The ATS 1 spacecraft shows its Syncom heritage, but at 351.5 kg (775 lb) on-orbit, with a length of 1.5 m (54 inches), was a larger and a much more capable satellite than its predecessor. In addition to a C-band transponder, ATS-1 carried a VHF experiment for communications with aircraft, ships, meteorological remote terminals, and to evaluate VHF satellite navigation. It used an eight-element phased array antenna with a 5 Watt transmitter for each element. The spacecraft was spin stabilized and had solar cells mounted around the barrel shaped body to provide 175 Watts of electrical power.

ATS-1 was the first satellite to use frequency division multiple access (FDMA) taking independently uplinked signals and converting them for downlink on a single carrier. This conserves uplink spectrum and also provides efficient power utilization on the downlink. ATS-1 also carried a black-and-white weather camera which transmitted the first full-disk Earth images from GEO. ATS-1 objectives were to test the experimental geostationary techniques of satellite orbit and motion, measure the orbital environment at 23,000 miles above the Earth’s surface, and transmit meteorological information (imagery and data) to surface ground stations.

The spacecraft was a cylinder 56 inches in diameter, 57 inches high and weighed 750 pounds. A phased array of eight whip antennas extended from the top, and a phased array of eight VHF antennas extended from the base. The sides of the cylinder were covered by 23,870 solar cells which, along with nicad batteries, provided the power for the craft. Two meteorological experiments were on board. One was a spin scan cloud camera which provided continuous, full-disk hemispheric images of the Sun-lit Earth every half hour. The spinning motion of the satellite generated line scans with a spatial resolution of 3.2 kilometers. This process took approximately 20 minutes for the full image, and then 10 minutes to reset the camera for a new image. The second experiment was Weather Facsimile (WEFAX), a data relay and re-transmission instrument. This instrument relayed data from the central ESSA data processing facility to APT ground stations located around the western hemisphere.

ATS-1 was placed in a transfer orbit directly over the equator over Ecuador. The transfer orbit meant that the satellite would drift slowly westward with time. The satellite eventually reached 151 degrees west (just east of Christmas Island) where it was deactivated on December 1,1978. Of the satellite’s 12-year life span, useful data was received for the first 6 (1966-1972).

One of the spectacular and promising events during 1966 was the successful operation of the ATS-I spin-scan camera. This camera is providing the first high-quality cloud-cover pictures taken from an equatorial synchronous satellite. These pictures show the disk of the Earth between 52 deg. Latitude N. And S. With a resolution approaching 3 km. The camera system can take the disk pictures once every 20 minutes and smaller areas more frequently, affording a potential continuous watch of global weather patterns. Another meteorological experiment on ATS-I transmits meteorological data (weather maps, cloud analyses, and spin-scan cloud pictures) from the ATS ground station at Mojave, California, via the satellite’s VHF transponder, to APT ground readout stations in the United States, Canada, Japan, Australia, and islands in the Pacific.

ATS-1 was launched December 7, 1966 and carried out an impressive array of experiments in communications and in collecting weather data. The communications hardware was still functioning nearly 20 years later when the spacecraft did not respond as intended to commands in 1985 and it slipped into a useless orbit.

ATS-1 Stats:

  • Launch Date:  December 7, 1966
  • Operational Period:   Deactivated on December 1, 1978
  • Launch Vehicle:    Atlas Agena D
  • Launch Site:    Cape Canaveral, FL
  • Type:    Weather Satellite

ATS- 2
Applications Technology Satellite-2 objectives were to test the experimental geostationary techniques of satellite orbit and motion; measure the orbital environment at 23,000 miles above the Earth’s surface; transmit meteorological information (imagery and data) to surface ground stations. The spacecraft failed to reach its intended orbit and was deactivated after the second stage of the Atlas rocket failed.
ATS-2 Stats:

  • Launch Date:    April 6, 1967
  • Operational Period:    April 6, 1967 failed
  • Launch Vehicle:    Atlas Agena D
  • Launch Site:   Cape Canaveral, FL
  • Type:  Weather Satellite

The objectives of the ATS-3 satellite was to develop the experimental geostationary techniques of satellite orbit and motion, measure the orbital environment at 23,000 miles above the Earth’s surface, and transmit meteorological information (imagery and data) to surface ground stations. Slightly larger than ATS-1, ATS-3 was a cylinder 60 inches in diameter, 72 inches high and weighed 805 pounds. A phased array of eight whip antennas extended from the top, and a phased array of eight VHF antennas extended from the base. The sides of the cylinder were covered by 23,870 solar cells which, along with nicad batteries, provided the power for the craft.

Three meteorological experiments were on board. One was a spin scan cloud camera which provided continuous, full-disk hemispheric images of the Sun-lit Earth every half hour. This camera was modified to produce color images. The spinning motion of the satellite generated line scans with a spatial resolution of 3.2 kilometers. This process took approximately 20 minutes for the full image, and then 10 minutes to reset the camera for a new image. The second experiment was an Image Dissector Camera which scanned the full-disk electronically rather than mechanically. The third experiment was Weather Facsimile (WEFAX), a data relay and re-transmission instrument. This instrument relayed data from the central ESSA data processing facility to APT ground stations located around the western hemisphere. In addition, images from the spin scan camera were also transmitted over WEFAX to APT stations.

ATS-3 was placed in a transfer orbit directly over the equator over 45 degrees west (ATS-1 was over the eastern Pacific at this time). The transfer orbit meant that the satellite would drift slowly westward with time. The satellite eventually reached 95 degrees west where it was deactivated with ATS-1 on December 1, 1978. Of the satellite’s 11-year life span, useful data was received for the first 8 (1967-1975).

ATS-3 Stats:

  • Launch Date:   November 5, 1967
  • Operational Period:    Deactivated with ATS-1 on December 1, 1978
  • Launch Vehicle:    Atlas Agena D
  • Launch Site:  Cape Canaveral, FL
  • Type:   Weather Satellite

The ATS-4 was launched to evaluate gravity-gradient stabilization and new imaging techniques for meteorological data retrieval. The spacecraft was deactivated after 61 minutes when the Centaur rocket failed to ignite. The satellite’s orbit was too low and the resulting atmospheric drag caused it to re-enter the Earth’s atmosphere.
Although the spacecraft was not stable, due to this low orbit, NASA engineers successfully turned on several of the experiments to collect as much information as possible before the satellite broke apart on October 17, 1968.

ATS-4 Stats:

  • Launch Date:  August 10, 1967
  • Operational Period:    Failed
  • Launch Vehicle:   Atlas Centaur
  • Launch Site:    Cape Canaveral, FL
  • Type:    Weather Satellite

The objective of the ATS-5 satellite was to evaluate gravity-gradient stabilization and new imaging techniques for meteorological data retrieval. Shortly after separation, the spacecraft began to wobble, and although control was restored and the satellite reached its destination (over the equator west of Ecuador), the satellite was spinning in the wrong direction. Although some of the instrumentation on board functioned properly despite the wrong spin, the gravitational gradient and imagery acquisition experiments failed. WEFAX was still operational although no images were transmitted.

ATS-5 Stats

  • Launch Date:    August 12, 1969
  • Operational Period:   Failed
  • Launch Vehicle:  Atlas Centaur
  • Launch Site:   Cape Canaveral, FL
  • Type:  Weather Satellite



The I-VI Reboost Program—Joan Tarro

WOW!  What an exciting program to have been lucky enough to have been the Program Office secretary on.

My career at Space & Communications was one of great joy.  To have been able to work with so many incredible people was a true gift.  To be honest, although I loved working on all the programs I worked on, I-VI Reboost was my favorite.

Dr. C. P. Rubin (Chuck) was the Program Manager and I felt honored that he asked me if I would like to work on that program with him (we had worked together previously).  I jumped at the opportunity.  It was a lot of hard work but so satisfying and fun.  I got to know all the astronauts and I was always kind of awe struck when I would call down to the Cape and ask to speak to one of the astronauts and they would come to the phone to talk to me (how many people can say they could do that).

I saved EVERYTHING regarding this program – all SCG Journals, newspaper articles, pictures, launch info, etc.

This program was the maiden voyage of the Endeavor (another exciting part of the program).  I was lucky enough to also have been able to see Endeavor return and land at Edward’s Air Force Base. To this day, it still makes me excited and smile when I think of working on this program.

Here is my list of some of the people that worked on the program:

Program Manager: Chuck Rubin,

Jason Endo, Asst. Prog. Mgr.

Len Dest, Mission Manager

System Engineering:

Don Patterson, LVI

Dennis Nell, Electronics

Willie Magee, Safety

Mark Altobelli

Product Assurance:

Dave Duckworth

S. Joe

Integration & Test/Launch Ops:

Ben Latuperissa

Ralph Yoak

Bill Amlotte

Bill Scanlon

Paul Ely

Structural Mechanics:

Sam Bassily, Manager

Bob Holman, Dynamics

Gunars Krumins

Bruce Linaman

Marygrace Literatus

Ray Ng

Thermal Control:

Stan Banash

Norm Pierson

Greg Reynolds

Design Integration:

Jason Endo, Manager

Al Byrne, Staff

Leo DelaCruz, Wire Harness

Dave Esposto, Structure Design

Dave Fox, Scheduler

Frank Freres, Manufacturing

Jeff Turney, GSE

Carl Jund, Super*Zip


Dave Husar

Carl Stammerjohn

Mass Properties:

Greg Johns


Phil Donatelli

John Ellison


Loren Slafer

Digital Electronics:

Russ Piombino, Manager

Mission Analysis Operations:

Ted Isaacs

Jerry Salvatore

Bruce Shuman


Intelsat VI F-3 Recovery


Editor’s Note:  I have combed my files and found some very interesting items on the Intelsat VI F-3 recovery.  The first piece is a Satellite Fact Sheet from the Boeing website, however I’m sure that it was originated by Hughes.  The other three items are NASA originated as indicated.  These convey much information about this recovery operation, however, the repsonsible Hughes engineers are not mentioned at all.  Anyone who has notes or recollections about our people that were involved in this operation are encouraged to comment on this post.

The most complex satellite rescue mission yet attempted was undertaken with the Intelsat VI F-3 reboost mission in 1992. It involved the first dual active rendezvous between a satellite and a space shuttle and the first in-orbit attachment of a solid propellant rocket motor.

As part of its maiden voyage, the space shuttle Endeavour rendezvoused with Intelsat VI  F-3, which was in a low Earth orbit since March 1990.  Astronauts maneuvered in space to capture the satellite, bring it into the shuttle bay, install the new motor, and then release it again into space. When Endeavour was moved a safe distance away, the new motor was fired to boost Intelsat VI F-3 into the geosynchronous orbit for which it was intended.


Intelsat VI F-3, one of five Hughes-built spacecraft in the Intelsat VI series, was launched on March 14, 1990. Because of a malfunction in the Titan 3 launch vehicle separation system, the second stage did not separate and prevented the satellite from continuing its ascent to its intended 22,300-mile geosynchronous orbit.

To prevent re-entry, satellite controllers had to react quickly to move the satellite to a safe orbit. Engineers in the INTELSAT Launch Control Center, working with telemetry, tracking, and command stations around the world, commanded the satellite to separate from its unused perigee kick motor (PKM), which took with it the burned-out Titan stage. Additional maneuvers, using the onboard liquid propellant, raised the orbit to 299 by 309 nautical miles to reduce the damaging effects of atomic oxygen on the silver solar cell interconnects that are integral to the spacecraft’s solar power generators. There the satellite remained, in a stable thermal and power state and in a safe but useless orbit, while engineers considered how best to raise F-3 to its appropriate orbit.

Hughes Aircraft Company’s Space and Communications Group built the spacecraft for the International Telecommunications Satellite Organization. The Intelsat VI series encompasses five satellites used to provide domestic and transoceanic voice, television, and data services to Earth stations in 180 countries. After the reboost mission, the F-3 spacecraft was positioned over the Atlantic Ocean to provide services between and among countries in North America, South America, Europe, and Africa.

Each Intelsat VI spacecraft has a capacity for three TV channels and 120,000 simultaneous two-way telephone circuits, using digital circuit multiplication equipment. Each also carries 48 active transponders, 38 at C-band and 10 at Ku-band.

 Mission Preparation

After studying several options, INTELSAT decided that it was more cost-effective and safer to bring a new motor to the satellite in space and reboost it from there, than to retrieve it, return it to ground, attach a new motor, and relaunch the satellite. Hughes was awarded a $43 million contract for special hardware, a new PKM, and mission support; and National Aeronautics and Space Administration (NASA) received a contract for a shuttle mission.

The new motor, built by United Technologies Corp. Chemical Systems Division, was an Orbus 21S solid rocket motor weighing 10,430 kilograms (23,000 pounds). It is called a PKM because the motor boosts the satellite from perigee (or lowest point in its elliptical orbit).

INTELSAT worked closely with Hughes and NASA in planning the recovery mission. Hughes built special hardware for the extravehicular activity and provided a specially designed cradle located in the aft payload bay to support the PKM and associated equipment during ascent in the shuttle.

Hughes also designed and built a docking adapter assembly. Astronauts will latch the adapter’s manually operated mounting clamps in space. The adapter will be released by pyrotechnically actuated bolt cutters by command from the ground after the motor is fired.

In providing mission support, Hughes engineers worked 25 feet underwater side by side with the astronauts who will walk in space to perform the reboost. Training was conducted underwater to simulate the effects of zero gravity, using NASA’s weightlessness environmental training facility in Houston, Texas.

 Operational Maneuvers

Rendezvous of Intelsat VI F-3 and the Endeavour involved intricate coordination of maneuvers of the two spacecraft. Immediately after shuttle launch, engineers in the INTELSAT Launch Control Center in Washington, D.C., and at stations around the world began sending a complex set of commands to the satellite to lower it from its 300-mile orbit. This was the first time that two spacecraft maneuvered simultaneously to achieve a rendezvous.

Intelsat VI F-3 and Endeavour rendezvoused within a “control box” volume that was defined by NASA approximately five hours after Endeavour’s takeoff. The control box extended over 6 degrees of arc in a roughly 200-nautical-mile circular orbit with an inclination of 28.35 degrees. Intelsat VI F-3 had to be in the control box within 46 hours after shuttle liftoff, and Endeavour completing the rendezvous 24 hours later.

About six hours before Endeavour’s approach, the satellite spun down via ground commands to its on-board thrusters and attitude control electronics. The spin speed was reduced to approximately 0.65 revolutions per minute. After the spindown was completed, the satellite was commanded into a safe configuration for rendezvous.

Endeavor, commanded by astronaut Dan Brandenstein, approached the satellite on flight day 4. Once the satellite and orbiter rendezvoused, Endeavour astronauts went outside the shuttle to capture the satellite, bring it into the shuttle bay, and attach the new perigee kick motor. The astronauts were expected to complete the tasks in 4.5 to 6 hours.

Extravehicular Activity

Approximately 90 minutes before the satellite was captured, two astronauts entered the payload bay. One astronaut was positioned on the manipulator foot restraint, which was attached to the remote manipulator system (RMS). After the RMS positioned the astronaut near the aft end of the satellite, he attached the capture bar to the satellite. Any satellite rotation was halted by the astronaut using a “steering wheel” in the center of the capture bar.

The astronauts then used the capture bar to dock the satellite to the shuttle. After the satellite was brought into the cargo bay, the two astronauts were positioned on diagonally opposite sides of the cradle. One attached an extension to the opposite end of the capture bar to help position the satellite within special alignment guides on the cradle and the docking adapter. These guides placed the satellite within range of the docking adapter clamps. The astronauts then latched the clamps.

Once the new motor was attached, the crew activated the power and the timer switches on each of the two staging and boost electronic units. The crew members then moved into the airlock prior to spacecraft deployment, which occurred at a time determined by INTELSAT. The shuttle crew in the cabin then fired the Super Zip to release the bay’s hold on the satellite. INTELSAT controllers established a command link with the satellite 35 minutes later, after it had reached a safe distance from the shuttle. The satellite’s control electronics were commanded on, its spin rate was increased, and it was placed in the proper attitude and prepared for motor ignition. The PKM was fired using a variable time delay that was commanded by the INTELSAT Launch Control Center.

The PKM thrusted the satellite into a supersynchronous elliptical transfer orbit with an apogee (or highest) altitude of about 45,000 miles (twice that of synchronous orbit). The higher transfer orbit altitude allowed a more efficient use of the PKM and reduced the amount of liquid propellant required to rotate the satellite’s orbital inclination toward Earth’s equatorial plane. By this means, the useful orbital lifetime of the satellite had been extended approximately 1.2 years to an estimated 10.8 years.

Over a period of several days, the controllers sent a series of commands to the satellite to take it to geostationary orbit. Once geostationary orbit was achieved, the controllers sent another series of commands to deploy the satellite’s solar drums and communications antennas.

After in-orbit testing, the satellite was expected to be in service over the Atlantic Ocean region by mid-1992.

NASA Press Release Sarah Loff May 13, 2015

On May 13, 1992, following the successful capture of the Intelsat VI satellite, three astronauts continue moving the 4.5 ton communications satellite into the space shuttle Endeavor’s cargo bay. A fellow crew member recorded this 70mm still frame from inside Endeavor’s cabin. Left to right, astronauts Richard J. Hieb, Thomas D. Akers and Pierre J. Thuot, cooperate on the effort to attach a specially designed grapple bar underneath the satellite. Thuot stands on the end of the Remote Manipulator System’s (RMS) arm while Hieb and Akers are on Portable Foot Restraints (PFR) affixed to Endeavor’s port side and the Multiple Support Structure (MPESS), respectively. The sections of Earth which form the backdrop for the scene are blanketed with thousands of square miles of clouds.EndeavorMissionThe Intelsat satellite, stranded in an unusable orbit since its launch aboard a Titan vehicle in March 1990, was equipped with a new perigee kick motor. The satellite was subsequently released into orbit and the new motor fired to put the spacecraft into geosynchronous orbit for operational use. The capture required three spacewalks: a planned one by astronaut Pierre J. Thuot and Richard J. Hieb who were unable to attach a capture bar to the satellite from a position on the RMS; a second unscheduled but identical attempt the following day; and finally an unscheduled but successful hand capture by Pierre J. Thuot and fellow crewmen Richard J. Hieb and Thomas D. Akers and Commander Daniel C. Brandenstein delicately maneuvered the orbiter to within a few feet of the 4.5 ton communications satellite.

The STS-49 mission, the first flight of shuttle Endeavor, set records for the first (and only, to date) spacewalk involving three astronauts; first shuttle mission to feature four spacewalks; first shuttle mission requiring three rendezvous with an orbiting spacecraft: first attachment of a live rocket motor to an orbiting satellite and first use of a drag chute during a shuttle landing.

NASA Photo Release 13 May 1992 Johnson Space Center

STS-49 ONBOARD SCENE—Florida’s Atlantic Coast and the Cape Canaveral area form the backdrop for this 70mm scene of Intelsat VI’s approach to the Space Shuttle Endeavor. Later, the seven-member crew was successful in capturikng the satellite and adding a perigee phase. The new motor allowed the needed boost for Intelsat, once the crewmembers had released it into space.IntelsatKSCNASA Photo Release 13 May 1992 Johnson Space Center

 STS-49 ONBOARD SCENE—This scene of Intelsat VI floating over a blue and white Earth greeted the STS-49 crewmembers as they prepared for their successful capture of the errant satellite. Within hours of this 35mm shot, three astronauts grabbed the satellite, moved it into the Endeavor’s cargo by, attached a perigee stage and released it into space. Later, a motor firing sent the vehicle on its way to a higher, geosynchronous orbit.IntelsatVIEarth

The following comment was provided by Dick (CR) Johnson:

Note that the failure of Intelsat VI F-3 to separate from the Titan L/V was due to a wiring error (not really a “malfunction”) in the separation system. The Titan P/L interface was configured to accommodate either a single passenger (lower P/L position) or dual passengers (lower and upper P/L positions) for injection and separation into a low earth orbit (LEO). The P/L separation system incorporated two sets of separation signal wiring – one for the upper P/L position and a second for the lower P/L position. It appears that a L/V technician connected the separation signal to the upper P/L position instead of the lower where VI F-3 was located. The physical separation interface was configured to separate the S/C, including the perigee kick motor (PKM), at the PKM interface. The only way to free the S/C from the spent Titan stage was to command PKM – S/C separation leaving the Intelsat VI F-3 in LEO with no PKM (not technically the L/V “second stage”) to inject it into Geosynchronous Transfer Orbit (GTO). (I believe that Intelsat brought legal action against Martin to recover the financial loss attributable to this wiring error.)