Approximately a decade following the Syncom initiative Hughes perceived that the major virtues of spacecraft spin stabilization would probably not be sufficient to continue Hughes’ dominance of the geosynchronous ComSat marketplace over the long term. Consequently an IR&D program was established to design and test a body stabilized ComSat “bus” variations of which were later offered in proposals for Intelsat V as well as for several government customers during the mid to late 1970’s. As it turned out, through a series of innovative design and marketplace initiatives, the reign of Hughes’ spin stabilized ComSats extended through the end of the 20th century, before being eclipsed by advanced body stabilized design technology.
Why Not Just Keep On Spinning?
The primary communication performance characteristics of a “high altitude” ComSat are “Effective Isotropic Radiated Power or EIRP (antenna gain times transmitted power directed toward the coverage area) and receiver “sensitivity” (antenna gain/receiver noise temperature or G/T). The most serious deficiency of an “all-spinning” ComSat design is the absence of a stable element for the mounting of highly directional (high gain) earth-oriented transmit/receive antenna beams1.
Additionally, the solar panel “prime power” available for powering communication transmitters and the power management of spacecraft “housekeeping” functions (telemetry, command, sensors, etc.) is derived from solar cells mounted on the vehicles cylindrical spinning “drum” rotor. Relative to a flat, sun-oriented panel, the cylindrical array “geometric” efficiency is about 32%. For a conventional spin stabilized design, the length (and, consequently the power output) of this cylindrical panel is limited by the requirement for passive dynamic stability that the inertial properties of the spacecraft be “disk shaped” with the spin axis moment of inertia larger than that of either transverse axis.
Both these antenna gain and prime power technical issues were largely mitigated by major design breakthroughs during the 1960’s and 1970’s. However, the spinning solar array “geometric inefficiency” could not be avoided. This fundamental prime power limitation prompted Hughes’ 1972 IR&D investment in a body stabilized S/C design program.
HS 361 IR&D Program
In 1972, Hughes executives recognized that future payload requirements would drive spacecraft design and push the limits of prime power because of the drum shaped solar array geometric inefficiency as well as the diameter and length constraints inherent in available launch vehicle fairing envelopes. While Hughes’ spin stabilized Comsat’s would continue to satisfy nearly all near-term mission requirements, it was understood that body stabilized spacecraft offered power advantages through the incorporation of large deployable, flat, sun-oriented solar arrays. Several satellites developed by our traditional competitors as well as the ATS-F and the Canadian Technology Satellite (CTS) had all demonstrated their capabilities in the early 70’s. Hughes’ challenge was to be ready to meet the growing future mission requirements of our commercial and government customers.
The decision to meet these future customer requirements with either a spin or body stabilized control system opened up new technical horizons. While many of Hughes’ existing technologies and capabilities were directly applicable to the implementation of body stabilized spacecraft designs, the hands on experience with body stabilized attitude control systems were limited to the days of the 1960’s Hughes Surveyor (“lunar soft lander”) spacecraft. The fundamental body stabilized design requirements were, in many areas, very different and not uniformly supported by Hughes’ prior design/manufacturing base. Establishing an internal hardware design/manufacturing capability with respect to spacecraft body stabilization was judged as crucial to positioning Hughes for future new business. The primary competition that lay ahead in 1975 was for the follow-on to Intelsat IV. As a result, a working group was set up in Dr. Leo Stoolman’s Systems organization and by the fall of 1972 the HS 361 Internal Research & Development (IR&D) project was put in place. Will Turk was named to lead the “HS 361” IR&D development program.
The key elements of the IR&D project were to:
a) Conduct tradeoffs and analyses for an Atlas Centaur Class Spacecraft design to gain insight into various nuances of the design including the unique transition from a spin stabilized transfer orbit into a body stabilized geosynchronous orbit deployment. Thor-Delta class missions were also considered during the later stages of the project.
b) To investigate alternative bus configurations to maximize payload performance while incorporating the most efficient power, thermal and propulsion systems (including an option for the incorporation of ion propulsion) and to evaluate deployable solar panel arrays and configurations including the Flexible Roll Up Solar Array (FRUSA) which was in development at Hughes and to test its deployment dynamics for this mission
c) To demonstrate substantive proof of concept of the attitude control system through a full up demonstration of the control systems acquisition and pointing control and to build/test a full scale engineering model of the primary body stabilized bus.
To these ends the Project achieved all of these goals in late 1974 and a HS 361 body stabilized spacecraft design and test program were completed incorporating an Intelsat V class payload based on early, postulated Intelsat communication design/performance requirements.
The 3800 pound HS 361 design incorporated a high speed momentum wheel gimbaled about two axes to control/stabilize spacecraft attitude to within 0.2 degree in pitch and roll and 0.5 degrees in yaw. Earth and sun sensors provide for attitude sensing along with inertial gyroscopes. Magnetic torque rods were employed to unload the wheel momentum accumulated from firing the propulsion jets and from external (primarily solar) torques. The design incorporated a solid propellant apogee kick motor to place the spacecraft into GEO following separation from the Centaur in a geosynchronous transfer orbit. Hydrazine thrusters were used to limit nutation divergence of the spin stabilized spacecraft during the transfer orbit and AKM burn phases to implement GEO N/S-E/W station-keeping. Prime power was derived from a sun oriented Flexible Roll Up Solar Array (FRUSA) which was selected for minimal solar panel weight. The FRUSA was partially deployed to provide power during the transfer orbit and subsequently stowed prior to the spacecraft’s erection/deployment in GEO. A maneuver to orient the spacecraft was performed and the momentum wheel was brought up to speed to stabilize/control the HS 361 in GEO. At this point the FRUSA solar array was fully deployed to provide up to 1 KW of power followed by deployment of the communication antennas. Ni-Cd batteries provided power during solar eclipses. The thermal control system incorporated mirrored thermal surfaces on the North and South faces of the spacecraft to reject internally generated heat loads (primarily from the communication payload’s high power transmitters).
For engineering and presentation purposes multiple models were built for testing purposes while several were created to demonstrate the key features of our design. A 1/3-scale model of the HS 361 was developed to show the design features and to demonstrate the various spacecraft states. The extension of the solar array during the transfer orbit and the on-orbit fully deployed solar panel and antennas could be shown with this single model and was used for both internal and customer presentations.
The engineering model, shown below, depicts the proposed structural elements containing the AKM thrust tube, propulsion system, the thermal control mirrors and the fully integrated communications payload.
While there are no pictures of the Attitude Control Lab built specifically to demonstrate our ability to design and build the HS-361 control system—this critical technology was incorporated in our design and the real time demonstration of the spacecraft attitude control was an IR&D project achievement.
The team that conducted the IR&D program is shown below against the background of the project’s hardware that also included a full- scale mockup of the basic HS 361 bus.
Late 1970’s Proposals
The HS 361 project enabled Hughes to position itself for several future commercial and government bids during the mid to late 1970’s. The first opportunity, the Intelsat V RFP, came along in 1975. The proposed Intelsat V body stabilized design was largely derived from the HS 361 development project. However, the Intelsat V design requirements led Proposal Manager Warren Nichols and Steve Pilcher (leading the technical design) to select several design and component alternates. The bus shape was modified and the tanks were positioned for improved inertial properties. A solid, hinge deployable, hard-backed solar panel was selected to meet higher power requirements.
In fact, Hughes offered two spacecraft designs in response to the Intelsat V RFP, a spin stabilized configuration based on the dual spin Intelsat IV A and as well as a body stabilized design. Although Hughes was not awarded the Intelsat V contract (nor any of the subsequent 1970’s body stabilized proposal opportunities), Intelsat confirmed that both of the Hughes S/C offerings were in full compliance with the Intelsat V’s technical requirements.
Another Decade of “Spinners”
The incorporation of “despun”, earth oriented antennas and the “Gyrostat”, dual spin configuration dramatically improved the communication performance of the Hughes’ family of spin stabilized spacecraft designs while retaining the major advantages of spin stabilization.2 These critical design initiatives enabled spin stabilized designs to maintain a competitive marketplace advantage throughout the 1970’s.
With the advent of NASA’s Space Transportation System (STS) or Space Shuttle initiative (1972), another opportunity for a major improvement in the communication performance and cost-effectiveness of spin stabilized ComSats became available. To attract satellite users and manufacturers to adopt the Shuttle as their launcher of choice, in the mid 1970’s NASA offered a very attractive price for a ride to low earth orbit (LEO). This price was a small fraction of the cost of launching on an expendable launch vehicle, but left the GEO S/C customer responsible for orbital transfer of their STS payload from LEO to GEO. As it turned out implementation of the required orbital transfer maneuvers was a great design match for the Hughes spin stabilized “integral propulsion” designs. Additionally the Shuttle’s payload bay envelope was much wider (15 feet in diameter) and longer (60 feet) than any available expendable launch vehicle’s.
Within a few months of NASA’s Shuttle launch pricing “formula” announcement, a spacecraft design effort lead by Dr. Harold Rosen produced a spin stabilized, “wide body” design optimized for, and launch- able only on, the Shuttle. This spacecraft design was dubbed “Syncom IV” and incorporated the spin stabilized integral propulsion (large solid rocket motor, or SRM, augmented by a liquid propulsion system) capable of the requisite orbital transfer from LEO to GEO. Since the Shuttle had not yet launched, The Syncom IV design was followed by the HS 376 spin stabilized configuration which was compatible with launch on either the Shuttle or the Delta expendable L/V. These S/C designs for Shuttle launch came to fruition years ahead of Hughes’ competitors and were unique in their ability to effect orbital transfer from LEO to GEO without the requirement for a separate upper stage. The Syncom IV “wide body” design concept was incorporated in the design of the “Leasat” ComSat for communication service leases to the US Navy. This Shuttle optimized “wide body” configuration was also purchased by the USAF and dubbed the “Multi Mission Bus” (MMB) which supported classified government communication payloads.3
During the first half of the 1980’s four Shuttle optimized wide body ”Leasat” ComSats for support of US Navy communications as well as numerous HS 376 domestic ComSats were launched into GEO by the Shuttle. Hughes enjoyed a major cost-effective competitive marketplace advantage employing Shuttle launches until the disastrous loss of the Shuttle Challenger and its crew in January of 1986. This watershed event prompted the government to preclude the additional STS missions which would have been necessary to support the growing demand for the launch of unmanned spacecraft4 so unfortunately, the Shuttle “party was over”!
Within weeks of the Challenger disaster, Hughes management realized that Hughes family of spin stabilized spacecraft was at a significant cost/performance disadvantage with launches restricted to expendable L/V’s. Consequently, the Space and Communications Group (S&CG) embarked on an ambitious design program, to develop a superior, state-of-the-art body stabilized spacecraft which became the HS 601 series.
This expenditure (in excess of $100 M) of internal funds to execute this major HS 601 design effort was approved by Hughes’ new owner, General Motors, and led by Ron Symmes with close design support from Dr. Rosen and Steve Pilcher. The HS 601 body stabilized spacecraft design was derived in part from S&CG’s 1970’s IR&D and proposal efforts, but was primarily aimed at “leap-fogging” the competition by incorporating the very latest in proven, high performance S/C ‘bus” space technology, such as the NiH batteries and bi-propellant propulsion adopted from the USAF’s MMB program. Most notably, since most available expendable launchers offered a lower cost ride into LEO (Vs GEO transfer orbit), the HS 601 design incorporated the Shuttle launched design legacy, as an option, for the transfer from LEO to GEO employing spin stabilized integral propulsion.5 The first HS 601, Australia’s Optus B1, was launched on the Chinese Long March 2E in August of 1992 from Xichang, China and is depicted in the artist’s rendering, below:
With the introduction of the very successful HS 601 and HS 702 body stabilized spacecraft “product lines” the era of Hughes’ ground-breaking spin stabilized ComSats effectively ended with the expiration of the 20th century.
1. See “Slicing the Bologna” subsection “Focusing” for the dramatic communication performance enhancements available employing “despun”, earth oriented spacecraft antennas.
2. See “Slicing the Bologna” subsection “Syncom” for the primary attributes of spacecraft spin stabilization.
3. The USAF’s “Multi-Mission Bus” introduced NiH battery technology to Hughes’ spacecraft implementing major improvements in battery efficiency, capacity and life. Also incorporated for the first time was mono-methyl hydrazine and nitrogen tetroxide (MMH/N2O4) higher Isp liquid bi-propellant propulsion.
4. The government granted two waivers to this prohibition based on the unavailability of expendable launch vehicles capable of accommodating Hughes’ “wide body” S/C configuration. STS launch of the fifth Leasat took place on 1/9/1990 preceded by the classified launch of the USAF’s MMB in the late 1980’s.
5. A more complete description of S&CG’s history with respect to their body stabilized design initiatives is presented in “Slicing the Bologna” under subsection “HS 601/702”.