Wobble Angle: 11 Apollo-derived Artificial-Gravity Space Stations (1963)

"Zero-G and I feel fine" — astronaut John Glenn in Earth orbit on board Mercury-Atlas 6 spacecraft Friendship 7, 20 February 1962. Image credit: NASA.
In early May 1963, Robert Mason and William Ferguson, engineers at the NASA Manned Spacecraft Center (MSC) in Houston, Texas, completed a study of 11 artificial-gravity Earth-orbital laboratory designs. Some might have argued that NASA engineers should have had had better things to do. After all, for two years NASA's main goal had been to land a man on the Moon and return him safely to the Earth before the Soviet Union did, and the U.S. program still lagged behind its Soviet counterpart.

At the time the MSC engineers completed their study, the U.S. record for weightless space endurance was held by John Glenn, the first American in orbit. During the Mercury-Atlas 6 mission (20 February 1962), he racked up a little less than five hours of weightless experience. On 24 May 1963, about two weeks after Mason and Ferguson completed their study, Scott Carpenter would match Glenn's feat during the Mercury-Atlas 7 mission.

The world record for weightless space endurance was, however, held by cosmonaut Andriyan Nikolayev, whose Vostok 3 spacecraft lifted off from Baikonur Cosmodrome on 11 August 1962. He orbited the Earth 64 times in 3 days, 22 hours, and 28 minutes, and landed on 15 August 1963. Apart from assurances that Nikolayev was in good health, the Soviet Union shared little information on his physical condition during or after his flight.

Lack of data on human responses to continuous weightlessness goes a long way toward explaining why NASA continued to study Earth-orbiting laboratories two years after beating the Russians to the Moon became a major U.S. goal. It seemed prudent to some to retain the option to launch a laboratory for studies of human health in weightlessness at least until astronauts could try to live in space for the period of time needed to accomplish an Apollo lunar landing mission.

Lack of data also explains why Mason and Ferguson studied artificial-gravity laboratory designs. If it were found that humans could not withstand weightlessness for extended periods, then it might become necessary to establish a lab in space where the human health effects and engineering requirements of spin-induced acceleration — which is what "artificial gravity" is — could be examined.

Before President John F. Kennedy put NASA on course for the lunar surface, an Earth-orbiting laboratory had been central to the agency's plans for the 1960s. By the end of 1962, the probable cost of the lunar program had begun to become clear. Grumbling had begun in Congress, placing pressure on President John F. Kennedy. The President in turn placed pressure on top NASA brass to contain space program costs.

It seemed possible that the Apollo lunar goal might be found wanting by either Kennedy or, if he lost his bid for reelection in November 1964, by his successor. If so, then NASA might wish to have handy a plan for an Earth-orbiting laboratory to serve as a replacement.

In all but one of their 11 designs, Mason and Ferguson had the laboratory and crew reach orbit together; the astronauts would ride in a modified Apollo Command and Service Module (CSM) spacecraft atop the laboratory's drum-shaped Mission Module (MM). CSM modifications included a much-shortened Service Module (SM) with only enough propulsive capability and supplies for the trip to the laboratory in 300-mile-high low-Earth orbit.

Mason and Ferguson focused their study on the extent of the shift in the station spin axis astronaut movement parallel to that axis would produce. They called that shift the "wobble angle."

This illustration from Mason and Ferguson's paper illustrates the "wobble angle." The line marked "Z" corresponds to the spin axis, which passes through the center of gravity of the orbiting laboratory. The Z at the top would, if the laboratory's spin remained entirely stable, always point directly at the Sun. Astronaut movement parallel to the Z line would, however, cause the spin axis to shift along the curving line labeled "Spin-axis trace." In this design, which corresponds to Laboratory Design 1 below, astronauts would need to contend with a wobble angle of about 43°. Mason and Ferguson likened this motion to "rolling of a ship."
The MSC engineers assumed that the orbiting laboratory MM (which would include any components specific to the artificial-gravity system), equipment, and the modified CSM would together weigh about 15 tons. Of that, five tons were allotted to the CSM. All of their designs retained the Saturn IB rocket second stage, the S-IVB, for use as a counterweight. Modified CSM, MM, and empty 10-ton S-IVB stage would orbit Earth at an altitude of 300 miles.

They set the spin rate at a maximum of four rotations per minute. At that rate, and at a distance of 40 feet from the spin axis, the acceleration the astronauts would feel would vary by 15% between their feet and their heads, with maximum acceleration being felt at their feet, farthest from the spin axis. Maximum acceleration was limited to one Earth gravity; minimum acceleration, to one lunar gravity (0.2 Earth gravities).

The 11 images that follow each include two views. The launch configuration is on the left and orbital configuration is on the right. In all but two, the Z axis/spin axis points at the viewer in both views; for Laboratory Designs 8 and 9, the Z axis in the launch configuration view is turned 90° relative to the orbital configuration view.

Laboratory Design 1: the first Mason and Ferguson artificial-gravity lab design is the simplest but also has one of the worst maximum wobble angles (about 43°). Crew couches in the CSM are at the minimum distance (40 feet) from the spin axis (Z); the entire two-deck, 7504-pound MM is too near the spin axis to avoid a greater than 15% variation in acceleration level between astronaut head and feet. Equipment weight is 12,496 pounds and pressurized volume is 2504 cubic feet. Thrusters on the laboratory would expend 52.9 pounds of propellant to spin the laboratory.
Laboratory design 2:  An alternative method of solar array deployment improves stability (wobble angle slightly more than 9°). MM structure weight is 8235 pounds and equipment weight is 11765 pounds. MM volume is 2505 cubic feet. Thrusters expend 55.3 pounds of propellants to spin up the laboratory. Unfortunately, no part of the CSM or MM is far enough from the spin axis to avoid a greater than 15% variation in acceleration level between astronaut head and feet.
Laboratory design 3:  Equipment modules of unspecified function deploy along the Y axis; this helps to reduce maximum wobble angle to about 3.5°. MM weight including the equipment modules is 12,492 pounds and equipment weight is 7508 pounds, the least of any of the designs. Thrusters expend 49.7 pounds of spin-up propellants; again, this is the least of any of the designs. MM volume is 2396 cubic feet. No part of the CSM or MM is far enough from the spin axis to avoid a greater than 15% variation in acceleration level between astronaut head and feet.
Laboratory design 4: A tunnel between the CSM and the MM places the CSM crew couches 43.3 feet from the spin axis. Unfortunately, the maximum distance from the spin axis within the MM is just 18.3 feet and the wobble angle is nearly 44°. MM structure weight is 8687 pounds and equipment weight is 11,313 pounds. Thrusters expend 50.7 pounds of propellants to spin up the laboratory. 
Laboratory design 5: The tunnel linking the CSM and MM is extendable, increasing the CSM crew couch/spin axis distance to 52.9 feet. The wobble angle is identical to that of design 4. MM structure weight is 8290 pounds and equipment weight is 11,710 pounds. Thrusters expend 65.7 pounds of propellants to spin up the laboratory. The 2400-cubic-foot MM entirely surrounds the spin axis; in theory, an astronaut at the spin axis would be weightless while the station spun around them. No part of the MM is far enough from the spin axis to avoid a greater than 15% variation in acceleration level between astronaut head and feet.
Laboratory design 6 includes a telescoping MM. The 45° wobble angle is the worst of the 11 designs. MM structure weight is 7505 pounds and equipment weight is 11,765 pounds. MM volume is just 1633 cubic feet, the least of any of the designs. Thrusters expend 68.3 pounds of propellants to spin up the laboratory. About half the MM is far enough from the spin axis to avoid a greater than 15% variation in acceleration level between astronaut head and feet.
Laboratory design 7 is similar to design 6, but its modified solar array configuration reduces its maximum wobble angle to slightly less than 29°. Structure weight is 7869 pounds, equipment weight is 12,131 pounds, and MM volume is 1743 cubic feet. Thrusters expend 68.3 pounds of propellants to spin up the laboratory.
Laboratory Design 8 combines features of designs 3 and 7 to achieve a wobble angle of slightly less than 2.5°. A new feature of this design is a docking port at the spin axis (Z) for a visiting CSM. Structure weight is 12,169 pounds, equipment weight is only 7831 pounds, and pressurized volume is 2048 cubic feet. Thrusters expend 66.9 pounds of propellants to spin up this design. All of the CSM and nearly all of the MM is far enough from the spin axis to avoid a greater than 15% variation in acceleration level between astronaut head and feet.
Laboratory Design 9 includes new structural elements: a "fork" and cables that permit the spent S-IVB stage to be pivoted 90° relative to its launch axis. This reduces the wobble angle to slightly less than 1° — the least of any of the 11 designs. Unfortunately, no part of the CSM or MM is far enough from the spin axis to avoid a greater than 15% variation in acceleration level between astronaut head and feet. Structure weigh is 8306 pounds and equipment weight, 11,694 pounds. Pressurized volume is 3118 cubic feet. Thrusters expend 64 pounds of propellants to spin up this design.
Laboratory Design 10 employs a "rigid support" and cables to pivot the spent S-IVB stage 90° relative to its launch axis. This reduces the wobble angle to 1°. The CSM crew couches and part of the MM are far enough from the spin axis to avoid a greater than 15% variation in acceleration level between astronaut head and feet. Structure weight is 8120 pounds and equipment weight is 11,880 pounds. Pressurized volume is 2400 cubic feet. Thrusters would expend 71.8 pounds of propellants to spin up this design.
Laboratory Design 11 includes no CSM in its launch configuration view because MM and equipment combined weight is too great; the CSM is displayed in the orbital configuration view as it would appear after it launched separately and docked with the MM in orbit. The pivoted S-IVB stage and solar panel configuration help to compensate for the large MM, yielding a wobble angle only slightly greater than design 9. Structure weight is 14,047 pounds and equipment weight is 15,953 pounds. Pressurized volume is by far the greatest of the 11 designs (3828 cubic feet), as is the amount of spin-up propellant required (116.1 pounds). All parts of CSM and MM are far enough from the spin axis to avoid a greater than 15% variation in acceleration level between astronaut head and feet. 

Project Apollo Conceptual Rotating Space Vehicle Designs Using Apollo Components for Simulation of Artificial Gravity, NASA Project Apollo Working Paper No. 1073, NASA Manned Spacecraft Center, 8 May 1963.

More Information

Space Station Resupply: The 1963 Plan to Turn the Apollo Spacecraft into a Space Freighter

To "G" or Not to "G" (1968)

"A True Gateway": Robert Gilruth's June 1968 Space Station Presentation

A Forgotten Rocket: The Saturn IB

Mars Airplane (1978)

Wings over Mars: the JPL Mars airplane swoops past a martian mountain so that its camera, mounted inside a clear plastic bubble on its belly, can turn sideways to image layers on the mountain slopes. Image credit: Jeff Bateman.
In the 1970s, as U.S. piloted spaceflight retreated to low-Earth orbit, NASA planning for advanced robotic Mars exploration missions came into its own. New information on the martian environment from Mariner 9 and the twin Vikings fueled engineer imaginations.

Many concepts that became actual missions in the 1990s and 2000s first received detailed study in the 1970s. Planners also looked at concepts that have yet to yield NASA missions: Mars sample return, balloons and blimps, small lander networks, and gliders and powered fixed-wing aircraft.

Spacecraft and space mission design and development decisions are complex and influenced by many factors. Scientific efficacy is but one factor considered when planning for a new exploration mission begins, and it is not always the most important one. Nevertheless, scientists are almost always involved at the outset even when they do not originate the mission concept under consideration. Often this involvement is achieved through the establishment of a supportive science working group for the proposed mission.

The Ad Hoc Mars Airplane Science Working Group met at the Jet Propulsion Laboratory (JPL) in Pasadena, California, on 8-9 May 1978, to review mission objectives and propose a possible Mars airplane instrument payload weighing between 40 and 100 kilograms. In its report, the Group noted that a Mars Airplane designed for landings and takeoffs would be able to collect samples in places other types of vehicles might find hard to reach. The plane might also be used to deploy small payloads at scattered locations by airdrop or landing.

Mostly, however, the Ad Hoc Science Working Group limited its deliberations to use of the plane as an aerial survey platform. The Group based its planning on a Mars airplane design derived from NASA Dryden Flight Research Center's "MiniSniffer" pilotless plane, which was designed to sample Earth's stratosphere.

The 300-kilogram airplane would arrive at Mars folded in an lozenge-shaped Viking-type aeroshell. After aeroshell parachute deployment and heat shield separation, it would spread its hinged wings to their full 21-meter span and detach from the parachute and aeroshell in mid-air.

Conceptual Mars airplane design. Image credit: Jeff Bateman.
Normally, the plane would cruise one kilometer above the martian surface, though it would be capable of flying as high as 7.5 kilometers. The 4.5-meter-diameter propeller at the front of its 6.35-meter-long fuselage would pull it through the thin (less than 1% of Earth atmosphere density) martian atmosphere at a speed of between 216 and 324 kilometers per hour.

Mars airplane endurance would depend on the weight of its payload and the choice of power plant. A plane with a 13-kilogram, 15-horsepower hydrazine-fueled piston motor, 187 kilograms of hydrazine fuel, and a 100-kilogram payload could, the Group estimated, fly up to 3000 kilometers in 7.5 hours, while one with a 20-kilogram electric motor, 180 kilograms of advanced lightweight batteries, and a 40-kilogram payload could fly up to 10,000 kilometers in 31 hours.

After it depleted its fuel or batteries, the plane would crash on Mars. The Group noted that the plane's short operational lifetime would dictate that its position after atmosphere entry be determined rapidly so that it could be directed quickly to its survey targets.

The Ad Hoc Group assumed that the Mars airplane would carry an inertial guidance system, radar and atmospheric-pressure altimeters, and terrain-following sensors (laser or radar) for navigation, and that these would serve double-duty as science instruments. The Group's selected science payload was intended to characterize possible landing sites for a follow-on Mars sample return mission and also to perform "topical" studies. The latter would address specific questions about Mars: for example, "Is Valles Marineris a rift valley?"

Visual imaging would be "fundamental" to the Mars airplane mission, so would receive top priority in the instrument suite. The Group determined that the airplane would be well-suited to serve as a camera platform because it would offer image resolution intermediate between orbiter and lander cameras and would obtain valuable "oblique" (from the side) images of the surface.

A Mars airplane might fly down a sinuous martian outflow channel, for example, collecting high-resolution images of layers exposed in its walls. The Mars airplane camera might be mounted on a movable platform inside a transparent dome on the plane's belly.

Other high-priority investigations would include wind speed, air pressure, and temperature measurements at various altitudes, infrared and gamma-ray spectroscopy and multispectral imaging to determine surface composition, and measurements of local magnetic fields. For magnetic field studies, the plane would fly a grid pattern over a selected region. The magnetometer, which might be mounted on a boom or a wingtip to minimize interference from airplane electrical sources, could also be used to seek out iron-rich surface materials and buried iron-rich volcanic structures.

The 1978 Mars airplane conceptual design effort fell victim to post-Viking disenchantment with Mars. Viking, which cost more than $1 billion in 1975 dollars, had been intended to find life, but its three biology experiments did not produce an unequivocally positive result. The Mars community did not at first recognize that it would need to restore support for Mars exploration before it proposed new Mars missions; that is, that Viking had made it more difficult to sell Mars exploration, not easier.

In addition, Space Shuttle development experienced setbacks. It was difficult to justify development of a vehicle for flying in the thin atmosphere of Mars when NASA had difficulty building one to fly in the thin upper atmosphere (and thicker lower atmosphere) of Earth.

Mars missions would resume, but not until 1992, when NASA launched a sophisticated orbiter called Mars Observer. The spacecraft was meant to inaugurate a new era of Mars exploration by providing a new overview of the planet. The loss of Mars Observer as it neared its destination on 25 September 1993 was a major setback; for a time, it appeared that recriminations over the very public failure might halt NASA Mars exploration.

The Curiosity rover landed in Gale Crater on 6 August 2012 and, after a checkout period, began its slow climb up the geologically complex layered slopes of Aeolus Mons (seen here in a color-corrected montage of images captured on 9 September 2015). At this writing, six-wheeled Curiosity has traveled about 22 kilometers. A Mars airplane could provide a perspective on Aeolus Mons, Valles Marineris, and other large features of Mars intermediate between that of a rover and that of an orbiter. Image credit: NASA.

Final Report of the Ad Hoc Mars Airplane Science Working Group, JPL Publication 78-89, NASA Jet Propulsion Laboratory, 1 November 1978.

Mars Airplane Presentation Material Presented at NASA Headquarters, JPL 760-198, Part II, Jet Propulsion Laboratory, 9 March 1978.

More Information

The Russians are Roving! The Russians are Roving! A 1970 JPL Plan for a 1979 Mars Rover

After Venus: Pioneer Mars Orbiter with Penetrators (1974)

Purple Pigeon: Mars Multi-Rover Mission (1977)

Prelude to Mars Sample Return: The Mars 1984 Mission (1977)

Making Propellants from Martian Air (1978)

To "G" Or Not to "G" (1968)

The quintessential space station: Wernher von Braun's revolving artificial-gravity station in Earth orbit. This classic 1952 painting by Chesley Bonestell, the Dean of Space Artists, includes near its hub a pill-shaped piloted space tug. In this view, the station might not be rotating; at least one of the two astronauts visible at center left is floating above its hull (perhaps they have just been tossed away by its spin). Though widely identified with von Braun, the spinning wheel station concept was first described in detail by Herman Poto─Źnik in 1928. Image credit: NASA.
Previously on this blog, I described the 1960s NASA push to make a large Earth-orbiting space station the "new Apollo" of the 1970s. I also discussed plans to exploit Apollo lunar program technology and techniques to conduct a low-cost post-Apollo piloted space program (the Apollo Applications Program, or AAP) that would include temporary space stations.

Both the proposed Space Station Program and AAP had looming over them a potentially crucial question: should NASA spin its future piloted spacecraft, in whole or in part, so that astronauts within could experience artificial gravity? During the longest piloted spaceflight of the era (Gemini VII, 4-18 December 1965), astronauts Frank Borman and James Lovell had orbited the Earth in weightlessness for nearly 14 days, clearing the way for Apollo lunar missions. Their flight encouraged AAP and station planners; it was widely recognized, however, that the meager biomedical results of a single two-week flight by two men in a cramped capsule could not be extrapolated to months-long stays on board a space station.

In a conversational memorandum dated 24 September 1968, E. Marion, an engineer with Bellcomm, NASA's Washington, DC-based planning contractor, examined whether space stations should be designed to provide artificial gravity or should assume that humans could adapt to weightlessness (which he called "abaria"). If the latter were true, then station complexity and cost might be greatly reduced.

Gemini 7 as viewed from Gemini 6, December 1965. Image credit: NASA.
Marion noted that the space medicine community tended to believe that astronauts could adapt to long-term abaria, but cautioned that this was "opinion, nothing more." "In other words," he explained, "it is possible that man can't physiologically adapt to long term abaria, but it is much more likely that he can."

He added that, even if sustained abaria were found to cause health problems, then spinning the entire station might not be necessary. The crew might get by with periodic sessions seated in a spinning centrifuge. Elastic bands in clothing could place limb and torso muscles under continuous tension and "lower body negative pressure boots" could give the heart a workout by pulling blood into the legs.

Marion wrote that artificial gravity might eliminate much astronaut training. Tools, furnishings, and equipment on board the artificial-gravity station — for example, "a plate of food" — could be identical to those used routinely on Earth. Training time reduction might, however, prove elusive; the artificial-gravity station would need to be "designed for abaric operation simply as a contingency" and its crew trained to use its backup abaric systems.

Marion speculated that space travelers might prefer abaria to artificial gravity. He wrote that astronauts — "a strikingly atypical population sample" — might, by virtue of their enthusiasm for new experiences, find that abaria would make "the long confinement of a space voyage" easier to stand. He suggested that, in the interest of astronaut behavioral health, missions might be planned to include both weightless and artificial-gravity periods.

The Bellcomm engineer wrote that astronauts performing work in abaria would probably be less "efficient" than those in artificial gravity — that "you can get more work out of an astronaut if you don't leave him weightless." Artificial gravity might thus enable "a smaller crew and a smaller station."

On the other hand, a major justification for the Space Station Program was the ability to perform experiments in weightlessness. Experiments might be designed to compensate for artificial gravity, Marion wrote, but at the cost of greater complexity and less efficiency. "It doesn't help to have an efficient astronaut running an inefficient experiment," he explained.

Experiments requiring abaria might be mounted in a central hub that would rotate against the station's spin direction to cancel out artificial gravity. Astronauts would enter the counter-rotating hub to operate the experiments. Marion noted, however, that the abaric hub might undercut "astronaut efficiency right when we need it the most — when he's working on the experiments."

Marion then offered three options for determining whether artificial gravity should be incorporated into the Space Station Program, each with "abaria OK" and "artificial-gravity required" alternatives, and provided cost estimates for all. He based these on AAP and Space Station Program schedules under consideration within NASA at the time he wrote his memorandum.

The schedule for AAP in September 1968 began with a mission on board a Workshop in Earth orbit in 1971. The AAP Workshop was called the "wet" Workshop because it would be launched with liquid propellants filling the volume the crew would inhabit in orbit.

AAP wet Workshop concept in 1967-1968. The docked Apollo Telescope Mount at upper left is based on the Apollo Lunar Module design. Image credit: NASA.
It would, in fact, be a modified S-IVB stage, the second stage of a two-stage Saturn IB rocket. The stage would include a long upper tank containing liquid hydrogen, a short lower tank for liquid oxygen, a J-2 rocket engine, and a special docking module bolted to the top of the liquid hydrogen tank. An Apollo Command and Service Module (CSM) spacecraft with a crew of three would ride into space atop the Saturn IB. The spacecraft would detach from the S-IVB second stage upon arrival in Earth orbit.

Controllers on the ground would then vent the S-IVB tanks and J-2 engine to clear them of residual propellants. The CSM would dock with the front (axial) port of the docking module, then its crew would fill the empty hydrogen tank with breathable air and move equipment and furnishings from the module into the tank to outfit it. They would live and work in abaria for 28 days, then would return to Earth.

A second CSM would reach the AAP Workshop at the end of 1971. The astronauts would reactivate it and live on board in abaria for 56 days. Soon after they returned to Earth, a third CSM, the last scheduled to visit the Workshop, would arrive bearing an Apollo Telescope Mount (ATM). The ATM would dock with a radial (side) port on the docking module and the CSM would dock with the axial port. The astronauts would use the ATM to study the Sun during their 56-day abaric mission.

The AAP plan included an option to launch a backup Workshop in mid-1972 if the 1971 Workshop failed. Alternately, the second Workshop might support a new series of missions if NASA received funding to expand AAP.

The Space Station Program artificial-gravity station design in Marion's September 1968 memorandum was barely described, but would probably have shared features with the two designs depicted in the NASA images above. The station at top would have reached Earth orbit atop a single Saturn V; the "million-pound" station at bottom would have required three Saturn V launches and orbital assembly. Both designs include a counter-reporting hub; an Apollo Command and Service Module (CSM) spacecraft is docked to the hub of the station at top.
NASA spacecraft development has generally followed a four-phase system, the details of which have varied considerably. Phase A, from which most proposed programs never emerge, encompasses preliminary analysis; at the time Marion wrote, the proposed Space Station Program was in Phase A. Phase B would see more detailed analysis and early design. During Phase C, detailed design and early manufacturing would take place. Phase D encompassed manufacturing and testing.

At the time Marion wrote, NASA planners anticipated that Space Station Program development Phase B might last six months in 1969. If so, then Phase C would last 18 months in 1970-1971, partially overlapping 42-month Phase D, which would begin in early 1971 and end in late 1974. The station would reach orbit in early 1975 and its first crew would arrive soon after.

The first of Marion's three artificial-gravity development options would assume that prolonged abaria would not pose a problem for station crews. AAP would not be used to confirm this assumption. The first crew would arrive on the station in mid-1975 for a prolonged stay in abaria. If they experienced adverse health effects, then a second crew might fly to confirm that these were caused by abaria.

If, based on their experience, it became clear that artificial gravity was necessary, NASA would halt the Space Station Program and spend two years designing, developing, and building a "G-kit" for attachment to a second station. Thus modified, the second station would reach orbit in early 1978.

Marion estimated that artificial-gravity development option 1 would cost just $700 million if the assumption that long-term abaria was acceptable turned out to be correct; this would make it the cheapest of all the alternatives. If artificial gravity were required, however, then delaying the program to modify the second station while keeping the NASA, contractor, researcher, and astronaut teams together would push total cost to $1.415 billion, making it the most expensive of all the alternatives.

Artificial-gravity development option 2 would see the Space Station Program postponed so that NASA could fly an abaric 120-day AAP mission using the backup Workshop in 1972-1973. Phase B would begin in late 1971, then Phase C would span 1972-1973. Toward the end of Phase C, station design would be finalized based on results of the long abaric AAP mission. Phase D would span from mid-1973 through the end of 1976. The station would reach orbit in 1977.

Marion estimated that artificial-gravity development option 2 would cost $900 million if abaria turned out to be acceptable. It would cost $1.015 billion if artificial gravity were required.

For artificial-gravity development option 3, the station would be built with part of its artificial-gravity hardware in place; specifically, it would include the counter-rotating hub as part of its basic structure. Phase A would begin in 1969, as in option 1, and NASA would launch the station in mid-1975.

At least one crew would then live on board in abaric conditions. If abaria were demonstrated to be acceptable, the Space Station Program could continue without artificial gravity (it might be added later as an experiment, if funds became available). If artificial gravity turned out to be necessary, then systems would be added to the orbiting station to complete its artificial-gravity configuration.

Though Marion did not say as much, it seems likely that artificial-gravity systems added to the station in late 1975-early 1976 would comprise a counterweight — probably a spent rocket stage — and cables or a truss for linking it to the station. The counterweight would be carefully positioned to place the counter-rotating hub at the station's spin center; this would ensure that it could provide an abaric environment for experiments. Astronauts would live on board the artificial-gravity station beginning in 1976.

Marion estimated that, if the Space Station Program continued without artificial gravity, then option 3 would cost $800 million. If artificial-gravity were required, then the cost would reach $915 million. He ended his memorandum by recommending that NASA choose option 3.


"To 'G' or not to 'G'," Bellcomm Memorandum for File, E. D. Marion, Bellcomm, Inc., 24 September 1968.

More Information

Space Station Gemini (1962)

Space Station Resupply: The 1963 Plan to Turn the Apollo Spacecraft Into a Space Freighter

Apollo Extension System Flight Mission Assignment Plan (1965)

"Without Hiatus": The Apollo Applications Program in June 1966

"Assuming that Everything Goes Perfectly Well in the Apollo Program. . ." (1967)

"A True Gateway": Robert Gilruth's June 1968 Space Station Presentation

McDonnell Douglas Phase B Space Station (1970)

An Alternate Station/Shuttle Evolution: The Spirit of '76 (1970)

Prelude to the Lunar Base Systems Study I: Lunar Oxygen (1983)

Climbing toward reusability: liftoff of the Space Shuttle Orbiter Columbia at the start of mission STS-2 (12-15 November 1981). Image credit: NASA.
At the heart of the Space Transportation System (STS) was the Space Shuttle. The first Space Shuttle mission, STS-1 (12-14 April 1981), was the first two-person Orbiter Flight Test (OFT) mission and the first flight of the Shuttle Orbiter Columbia. The second OFT, STS-2 (12-15 November 1981), had to be cut to two days in orbit from a planned five following the failure of one of Columbia's three electricity-producing fuel cells. Nevertheless, STS-2, the first reflight of a reusable spacecraft and the first flight of the Canada-built Remote Manipulator System (RMS) robot arm, was viewed as a success.

When Columbia glided to a landing for the second time, the form the STS would eventually take was still poorly defined. It would remain so at least until the destruction of the Shuttle Orbiter Challenger (28 January 1986) at the start of the STS-51L, the 25th flight of the Shuttle Program. The loss of Challenger and her seven-member crew marked the end of the optimistic first phase of the Space Shuttle Program.

Before that, however, the STS seemed ripe for augmentation. It would, of course, include expendable rocket stages attached to satellites carried to low-Earth orbit (LEO) in the Shuttle Orbiter Payload Bay; these "upper stages" were meant mainly to boost payloads from LEO to geosynchronous orbit (GEO), but could also launch robotic spacecraft from LEO on interplanetary trajectories. In addition, the STS would include Spacelab, a European-built system of Payload Bay-mounted laboratory modules and scientific instrument pallets. Development of upper stages and Spacelab had commenced in the 1970s, shortly after Space Shuttle development began.

Many saw the stages and Spacelab as interim steps toward more complex and competent STS elements. The former, it was expected, would lead to a reusable Orbital Transfer Vehicle (OTV); the latter, to a Space Station assembled in LEO from Orbiter-launched components. The OTV and its "little brother," the Orbital Maneuvering Vehicle (OMV), were typically seen as auxiliary vehicles based at the Space Station.

NASA Johnson Space Center (JSC) in Houston, Texas, anticipated that the Space Station would support ambitious space construction projects; for example, large communications platforms in GEO. In 1979, inspired in part by its involvement in joint Department of Energy/NASA Solar Power Satellite studies, JSC studied a Space Station concept it called the Space Operations Center (SOC). After an initial flurry of planning activity, Shuttle delays put the SOC on the back burner; immediately after STS-1, however, JSC efforts to promote the assembly base in LEO kicked into high gear.

The Space Operation Center (SOC), a Shuttle-launched low-Earth orbit (LEO) space station, as it was envisioned in 1982. Intended as an assembly and repair base, the SOC would have included a "surrogate" Shuttle Orbiter Payload Bay (lower center) for satellite servicing and a hexagonal hangar (center right) for storing and servicing the Orbital Transfer Vehicle (OTV) and Orbital Maneuvering Vehicle (OMV). Image credit: NASA. 
A remote-controlled Orbital Maneuvering Vehicle (OMV), a small "space tug," closes in on the Gamma Ray Observatory (GRO) in this illustration from 1986. After its launch in 1991, GRO was renamed the Compton GRO. The reusable modular OMV, intended as an auxiliary vehicle for extending Shuttle Orbiter and Space Station capabilities, was cancelled in 1987 in the aftermath of the Challenger accident. Image credit: TRW/NASA.
JSC is widely known as the home base of the astronauts and site of the Mission Control Center. Less well known, perhaps, is its long-time contribution to lunar and planetary science. The Lunar Receiving Laboratory (LRL), built in 1967, was used for analysis and storage of lunar geologic samples beginning with Apollo 11 in July 1969. JSC also played a key role in organizing the annual Lunar Science Conference (LSC), the first of which was held in Houston in January 1970.

At the time, JSC was called the Manned Spacecraft Center (MSC). It was renamed in 1973 after the death of President Lyndon Baines Johnson. The LSC was renamed the Lunar and Planetary Science Conference (LPSC) in 1978.

Michael Duke was on hand when the Apollo 11 samples arrived at the LRL; he was Lunar Sample Principal Investigator for the mission. In July 1969, he had been working for the U.S. Geological Survey (USGS) Branch of Astrogeology for six years. In 1970, he left USGS to become Lunar Sample Curator at MSC, a post he held until 1977, when he became Chief of the JSC Planetary and Earth Sciences Division.

Shortly after STS-2, Duke and another geologist in his division, Wendell Mendell, became concerned that developing the SOC and other proposed STS elements might mean reduced funding for NASA science programs. Space science at NASA was already hurting; the new Administration of President Ronald Reagan had made deep cuts. Rather than oppose new STS elements, Duke and Mendell sought ways that the SOC, OTV, and other proposed hardware could advance scientific exploration. Specifically, they sought to make the case for a base on the Moon.

Their efforts in some ways paralleled those of lunar scientists at the dawn of the Apollo Program, when lunar science barely existed as a field of study. Much like those early pioneers, Duke and Mendell sought to find and bring together individuals and organizations to build a constituency. Initially, they found prospective lunar base allies through informal, low-profile contacts. By late 1982, however, it was time to go public.

This they did by organizing three public special sessions at the 14th LPSC, which was held at NASA JSC in March 1983. The lunar sessions were titled "Return to the Moon" and "Future Lunar Programs." The third session, "Prospects for Planetary Exploration," sought to tie their lunar base efforts to the interests of the broader Solar System exploration community.

In their introduction to the lunar special sessions abstract volume, Duke and Mendell explained that "very little vision is required to see the [STS] reaching to the Moon." They argued that "the lunar option requires decisions today — but not dramatic ones." They pitched a Fiscal Year (FY) 1985 start for the lunar base program, but took pains to stress that the lunar base would need little or no new dedicated funding before FY 1991 or FY 1992.

In fact, they expected that lunar capability would grow more or less naturally from the STS in the late 1990s, several years after the SOC, OTV, and OMV were in place to support GEO missions. The amount of energy required to put a satellite into GEO is, after all, very nearly the same as that needed to put a payload into low-lunar orbit (LLO).

The lunar special sessions abstract volume included 22 abstracts by more than 30 authors and co-authors. The abstracts covered topics ranging from philosophy, law, and economics to geology, physiology, and energy. Of particular interest was an abstract by Hubert Davis, Senior Vice President of Houston-based Eagle Engineering, Incorporated (EEI). In it, he proposed to extend the STS to the Moon. His aim: to make the STS more economical by mining, refining, and using as rocket propellant oxygen chemically locked in lunar dirt.

EEI had been established by retired NASA JSC engineers. Davis, a co-founder of the company, began his career as an aircraft maintenance engineer in the U.S. Air Force during the Korean War. Inspired by President John F. Kennedy's May 1961 call for a man on the Moon, he went to work for the MSC Power & Propulsion Division in March 1962. He oversaw the test program for Lunar Module 5; better known as Eagle, it bore Apollo 11 astronauts Neil Armstrong and Edwin "Buzz" Aldrin to the lunar surface on 20 July 1969.

As Apollo ended, Davis transferred to the MSC Special Projects Office, where he studied new STS elements — cheap solid-propellant STS upper stages and a Shuttle-derived heavy-lift launcher — as well as Solar Power Satellites manufactured from lunar materials. He took early retirement in 1978 after being made the JSC Engineering Directorate representative to the Space Shuttle Program Office; in a 2009 NASA JSC oral history interview, Davis explained that he had left the space agency at age 48 because he found the Shuttle oversight job to be boring.

The industry magazine Aviation Week & Space Technology made his presentation the centerpiece of its coverage of the LPSC lunar base special sessions. Two months following the special sessions Davis published an expanded version of his abstract and presentation as an EEI report called Lunar Oxygen Impact Upon STS Effectiveness.

Davis acknowledged that his study was incomplete and preliminary. He did not, for example, examine how his oxygen mining and refining facility would be established on the Moon. That the study was preliminary explained why it was incomplete; if it could not provide an early indication that lunar oxygen would enhance STS capabilities, Davis argued, then there would be little point in conducting a more detailed study of the concept.

To illustrate how lunar-produced liquid oxygen might enhance the STS, Davis used as an example STS-40, which was scheduled to place the Galileo spacecraft into LEO on 30 May 1986. Galileo would reach LEO attached to an expendable Centaur G' upper stage which, following release from the Orbiter payload bay, would launch it out of LEO on a direct trajectory to Jupiter.

The Galileo spacecraft was expected to weigh 2510 kilograms and the Centaur G' without propellants, 2650 kilograms. Support equipment for maintaining Galileo and Centaur G' in the payload bay would weigh 470 kilograms and 3640 kilograms, respectively. Filling the Centaur G' large tank with low-density liquid hydrogen fuel would add 3310 kilograms to STS-40's payload weight; filling the Centaur G' small tank with dense liquid oxygen oxidizer would increase payload weight by a whopping 16,570 kilograms. STS-40 payload weight thus came to 29,480 kilograms, with liquid oxygen making up 56% of the total.

The Galileo Jupiter spacecraft and an expendable Centaur G' upper stage move away from the Shuttle Orbiter that deployed them in low-Earth orbit. The large-diameter forward section of the stage (center), to which Galileo is attached, contains low-density liquid hydrogen fuel; the small-diameter aft section, to which are attached two rocket motors, high-density liquid oxygen oxidizer. Image credit: NASA.
Davis then imagined a 2510-kilogram Shuttle-launched payload in the first decade of the 21st century, after the STS had been extended to the Moon. A reusable OTV would replace the Centaur G'. Permanently basing the OTV at the SOC meant that it would not add to Shuttle payload weight.

The Shuttle Orbiter would carry a 660-kilogram tank containing 3310 kilograms of liquid hydrogen for fueling the OTV in LEO. The propellant dump in LEO near the SOC would provide the OTV with lunar liquid oxygen. Support equipment in the payload bay for the spacecraft and hydrogen tank would bring the total Shuttle payload weight to just 7280 kilograms, or about one quarter of the STS-40 total.

A chemical process called hydrogen reduction of ilmenite formed the basis of the lunar oxygen STS infrastructure. Ilmenite (chemical formula FeTiO3), a titanium ore, is a mineral common in the basaltic rocks, dirt, and dust that form the dark-hued lunar plains known as maria (Latin for "seas").

Davis focused on ilmenite rather than lunar polar ice — which can provide both hydrogen and oxygen — because in 1983 no one knew that water ice exists at the lunar poles. Though the lunar polar ice hypothesis was by then more than 20 years old, the first evidence that it might be correct would not be found until 1994, when Clementine became the first spacecraft to explore the Moon from lunar polar orbit.

Mining robots would continuously gather ilmenite-rich lunar dirt at a rate of 28 metric tons per hour and deliver it to a separation facility. The dirt would first be sieved to remove large dirt particles, clods, and rocks. The resulting fine-grained dirt and dust would then be heated and subjected to an electrostatic process that would separate the ilmenite at a rate of 2.27 metric tons per hour.

The ilmenite would be moved to the hydrogen-reduction unit, where it would be exposed to hydrogen gas at 700° Celsius (C) and 2.7 Earth atmospheres of pressure. The hydrogen would bind with and free the oxygen bonded to the iron. This would yield water vapor at a rate of 0.26 metric tons per hour, which would be cooled, condensed, and subjected to electrolysis, splitting it into oxygen and hydrogen.

The oxygen would be chilled until it condensed into dense liquid, then stored in spherical tanks. The hydrogen would be returned to the reduction unit for reuse and the powdery titanium oxide and iron left over from the reduction process — about 90% of the original mass of the lunar dirt — would be stacked out of the way for possible future use.

The facility would use a little more than six megawatts of electricity continuously; this might be reduced if waste heat from the refining process could be exploited effectively. Davis estimated that his mining and refining facility could produce 150 metric tons of liquid oxygen per month.

The Aft Cargo Carrier (ACC) — the blue short cylinder and truncated cone attached at left to the bottom of the orange Space Shuttle External Tank (ET) — would augment the 4.57-by-18.28-meter Shuttle Orbiter Payload Bay (the blue cylinder within the Orbiter outlined by dashed lines). The 9.72-meter-long, 8.38-meter-wide ACC would, among other things, facilitate launch of OTV components and spherical tanks filled with liquid hydrogen. Image credit: Martin Marietta.
This painting by Eagle Engineering artist Pat Rawlings displays elements of Davis's proposed lunar oxygen STS infrastructure in LEO. The close proximity of the elements is schematic, not realistic. At lower right, a remote-controlled OMV detaches a spherical tank filled with liquid hydrogen from an ET/ACC. The LEO propellant dump is at lower left. Small in the distance above it, silhouetted against the Earth, is a Shuttle Orbiter. A version of the SOC is depicted to the right of the Orbiter. In the foreground, a small OTV prepares to leave Earth orbit; liquid hydrogen tanks bound for lunar orbit ride on its bowl-shaped rigid aerobrake heat shield.
Davis described his lunar oxygen STS infrastructure in operation. A Space Shuttle would launch liquid hydrogen for the LEO propellant dump in a spherical tank inside an Aft Cargo Carrier (ACC) fitted over the dome-shaped end of its External Tank (ET).

Normally, the ET would separate from the Orbiter as it neared orbital velocity, tumble and break up in the upper atmosphere, and fall into the Indian Ocean. When the ACC was attached, however, the Shuttle Orbiter would boost the ET/ACC combination to LEO and separate. A remote-controlled OMV based at the SOC would then detach the hydrogen tank from the ACC and move it to the propellant dump, where refrigeration and high-tech insulation would ensure that no hydrogen was lost to boil-off.

Zero liquid hydrogen boil-off was, Davis explained, critical to making his lunar oxygen STS infrastructure viable. He wrote that the Centaur G' stage was expected to lose about 3% of its liquid hydrogen to boil-off per day. A similar boil-off rate at any point in his lunar oxygen STS infrastructure would be "intolerable."

Davis assumed two types of modular OTVs, each of which could be tailored to carry out several types of missions. The OTVs, clusters of spherical propellant tanks linked by struts, would perform roundtrip missions between LEO and GEO and between LEO and an LLO SOC and propellant dump. The OTVs could operate with or without a pressurized module containing a crew.

The smaller OTV, which would burn 25 metric tons of liquid hydrogen and liquid oxygen during a voyage from LEO to LLO and back, would include a rigid aerobrake heat shield 18 meters wide. The heat shield, which would include thermal protection tiles akin to those attached to the Space Shuttle Orbiter, would enable the OTV to use atmospheric drag to capture into LEO with minimal propellant expenditure. The smaller OTV could transport nearly 43 metric tons of liquid hydrogen from the LEO propellant dump to its counterpart in LLO.

A single-stage Lunar Module lander based on the smaller OTV design would burn 28 metric tons of liquid hydrogen/liquid oxygen propellants to travel from LLO to the lunar surface and back. It would be capable of lifting 41 metric tons of liquid oxygen from the lunar surface to the LLO propellant dump.

Davis used Lunar Module landing gear as an example of how hardware in his lunar oxygen STS infrastructure would need to be optimized to reduce mass. The Apollo Lunar Module's four landing legs and foot pads accounted for 3.3% of its landed weight; the small OTV-based Lunar Module would exploit new materials and improved understanding of the lunar surface to reduce the figure to 2%.

On the Moon: at lower left, a robotic front-end loader scoops lunar dirt; behind it, another deposits dirt in a hopper at the start of the lunar oxygen refining process. Cables strung on poles link the lunar oxygen refinery to a nuclear reactor just over the horizon. Lunar oxygen is liquified and poured into a tank at center right; the filled tank will be added to the stack of tanks in the background at center top. Conveyor belts transport tailings to a storage area at upper left, just beyond the Lunar Module launch & landing pad, and to the open pit mine at center left. Image credit: Pat Rawlings/Eagle Engineering, Incorporated/NASA.
In lunar orbit: in the foreground, a large OTV with a stowed white ballute heat shield prepares to depart LLO for Earth orbit carrying a cargo of lunar oxygen. In the background, the LLO propellant dump orbits close by the lunar SOC. Meanwhile, a Lunar Module bearing lunar liquid oxygen moves in to dock with the propellant dump. Image credit: Pat Rawlings/Eagle Engineering, Incorporated/NASA.
In its basic form, the larger OTV would carry a propellant load of 33 metric tons. Two such OTVs could be combined to form an OTV with a propellant load of 78 tons. The latter configuration would be capable of transporting more than 200 metric tons of lunar oxygen from LLO to the LEO propellant dump. This would, he calculated, require an aerobrake heat shield about 115 meters wide; that is, wider than an American football field with end zones.

One might be forgiven for asking why such a large lunar oxygen cargo was necessary; that is, why Davis did not propose transporting it to LEO using several smaller OTVs spaced out over time. He explained that minimum-energy opportunities for travel from the LLO propellant dump to LEO would occur less than once per month. They would be infrequent because the OTV could only depart LLO as its orbital plane coincided with that of the LEO propellant dump. To do otherwise would demand plane-change maneuvers that would contribute toward making the lunar oxygen STS infrastructure uneconomical.

During aerobraking, the OTV would pass through Earth's upper atmosphere at an altitude of between 50 and 100 kilometers so that atmospheric drag could reduce its speed. The OTV would then climb back into space toward an apogee (orbit high point) near SOC altitude (about 400 kilometers). At apogee, it would ignite its engines to raise its perigee (orbit low point) out of Earth's atmosphere. For the perigee-raise maneuver, Davis budgeted only enough propellants to change OTV speed by 100 meters per second. He suggested that, if further analysis showed this to be insufficient, then an SOC-based OMV might retrieve the OTV and lunar oxygen payload at apogee.

In neither the small OTV nor the large OTV case could aerobrake heat shield mass exceed 3.5% of OTV mass at the time of Earth atmosphere entry. Davis focused on heat shield mass reduction because other OTV systems were already optimized, OTV propellants had been trimmed to the bare minimum required, and reducing the liquid oxygen cargo would defeat the purpose of the exercise. He conceded that cutting aerobrake heat shield mass so dramatically might constitute a significant technical challenge; most OTV studies, he explained, had assumed that the heat shield would make up at least 10% of OTV mass at Earth atmosphere entry.

Ballute in action. This is representative of ballutes in general; the inflatable heat shield is shown here attached to a single-engine cylindrical tug, not the two-engine large OTV Davis described, and aerobraking events take place at higher altitudes than in his scenario. Image credit: NASA.
To reduce the mass of the large OTV heat shield, Davis suggested that it might take the form of an expendable fabric ballute ("balloon-parachute"). The OTV would point its twin engines in its direction of motion as the donut-shaped ballute inflated; the engines would then operate in "idle" mode to create a relatively cool gas barrier between the ballute surface and the high-temperature plasma generated in front of the ballute by Earth atmosphere reentry at lunar-return speed (3.2 kilometers per second).

Davis used computer models to attempt to determine the Mass Payback Ratio (MPR) of his proposed lunar oxygen STS infrastructure. An MPR of 1 would mean that the mass of resources (mainly propellants) expended to exploit lunar oxygen would equal the mass of the lunar oxygen supplied to LEO. NASA would thus gain nothing from putting lunar oxygen to work in the STS. If, on the other hand, the mass of the lunar oxygen delivered to LEO exceeded the mass of the resources needed to exploit it, then more detailed study might be justified.

Davis cited a computer model that included 25 roundtrip OTV flights between LEO and LLO and 103 roundtrip Lunar Module flights between LLO and the lunar surface. He wrote that, in exchange for 983 metric tons of liquid hydrogen, hydrogen tanks, and OTV attitude-control system propellant dispatched to the Moon, 2414 metric tons of lunar liquid oxygen would arrive in the LEO. He judged that this quantity could support 90 OTV flights between LEO and GEO over a period of about five years.

This indicated a preliminary MPR of 2.45, which, Davis wrote, justified additional study. He anticipated, however, that it probably would not provide enough margin to maintain a positive MPR if the mass of hardware and propellants required to establish and maintain the lunar oxygen STS infrastructure were taken into consideration.

Davis did not provide weight estimates for the LEO propellant dump, the LLO propellant dump and LLO SOC, and the Lunar Modules. Neither did he estimate the weight of the Earth-launched liquid hydrogen and liquid oxygen propellants needed to initiate the lunar oxygen STS infrastructure, nor the weight of Earth-launched liquid hydrogen needed to fly resupply and crew rotation missions after lunar oxygen became available. He assumed that the OTVs and LEO SOC would be built for LEO and GEO operations even if NASA did not return to the Moon, so disregarded their weight in his model.

He did, however, provide a weight estimate for the lunar surface mining and refining facility. Mining robots, a habitat for housing 10 facility caretakers, refining equipment, storage tanks, a nuclear reactor for generating electricity, radiator panels, and other equipment would have a combined weight of 437 metric tons. Adding this to the 983 tons of hydrogen, tanks, and attitude-control propellant would lead to an MPR of only 1.7.

If, somehow, the MPR remained sufficiently favorable after more detailed technical studies yielded credible complete weight estimates, then complex economic analyses would follow. These would, Davis explained, be based on real-world dollars and would take into account societal factors such as "affordability."

Davis conceded that extending the STS to the Moon probably could not be justified solely on the basis of economics. He argued that lunar resources should nevertheless be developed. He cited a January 1982 Los Alamos National Laboratory (LANL) proposal for an international research laboratory on the Moon; it promised wide-ranging scientific, economic, political, and defense benefits. With a nod to the political language of the 1980s United States, Davis declared that the "vitality of Free World commerce and physical security would be greatly increased by the presence of. . .resources in space."

This post is the first in a series on lunar base planning in the 1980s centered on activities at NASA JSC. The next installment will examine NASA JSC's March 1984 in-house Lunar Surface Return study.


Space Operations Center presentation materials, NASA Johnson Space Center, 18 January 1982.

"NASA Conference to Highlight Return to the Moon," NASA News Release 83-007, Steve Nesbitt, no date (March 1983).

"Economic Benefits of Lunar Base Cited," E. Bulban, Aviation Week & Space Technology, 18 April 1983, pp. 132-133, 135-137.

Fourteenth Lunar and Planetary Science Conference Special Sessions Abstracts — Return to the Moon — March 16, 1983, Future Lunar Program — March 17, 1983, LPI Contribution 500, Lunar and Planetary Institute, 1983.

Lunar Oxygen Impact Upon STS Effectiveness, Report No. 8363, Hubert Davis, Eagle Engineering, Inc., May 1983.

"Return to the Moon," Andrew Chaikin, Sky & Telescope, June 1983, p. 493.

NASA Johnson Space Center Oral History Project Edited Oral History Transcript: Hubert P. Davis, 28 July 2009 (https://historycollection.jsc.nasa.gov/JSCHistoryPortal/history/oral_histories/DavisHP/DavisHP_7-28-09.htm — accessed 20 March 2020).

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"A Vision of the Future": Military Uses of the Moon and Asteroids (1983)

Mission to the Mantle: Michael Duke's Moonrise (1999-2009)

Apollo Applications Program: Lunar Module Relay Experiment Laboratory (1966)

Artist's impression of Syncom 2. Image credit: NASA.
In October 1945, space advocate and author Arthur C. Clarke published a bold proposal in the pages of the popular British radio magazine Wireless World. He first explained that the speed of an object orbiting 35,786 kilometers above Earth's equator would match — that is, be synchronous with — Earth's equatorial rotation speed. From the perspective of people on Earth, such an object would appear to hover over one spot on the equator.

Clarke then proposed a network of three such Geosynchronous Earth Orbit (GEO) satellites spaced equidistantly along the equator. These would, he wrote, be well placed to relay radio signals all over the world.

By most accounts, his proposal was not taken seriously. German V-2 missiles had demonstrated that large rockets needed to launch satellites were possible; however, most people with an interest in radio communication saw his GEO radio-relay network as a project for the far future. Everyone else, of course, paid no attention whatsoever.

Less than 20 years later (26 July 1963), NASA launched drum-shaped Syncom 2 (image at top of post). Through a series of careful maneuvers, the Hughes Aircraft Company-built satellite reached a 35,786-kilometer-high orbit inclined 33° relative to Earth's equator on 16 August 1963.

In its inclined synchronous orbit, 68-kilogram Syncom 2 oscillated daily along a 66°-long path centered over a spot on the equator at 55° west longitude. This took it over the North Atlantic and Brazil, enabling test transmissions between North America, Europe, South America, and Africa.

A year later (19 August 1964), NASA launched Syncom 3 to a point directly over the equator, making it the world's first geostationary comsat. From its mid-Pacific position at the intersection of the equator and the International Date Line, Syncom 3 was well placed to relay TV signals from the 1964 Tokyo Olympic Games to North America.

On 6 April 1965, NASA launched Intelsat I, the first commercial GEO comsat. Nicknamed "Early Bird" by the international consortium that funded it, Syncom-derived Intelsat I remained operational until January 1969. It was turned on again briefly in July 1969 to relay signals from Apollo 11, the first piloted lunar landing mission.

A year after the Intelsat I launch, Samuel Fordyce of the NASA Headquarters Office of Manned Space Flight (OMSF) circulated a memorandum in which he proposed that an Apollo Lunar Module (LM) be outfitted as a radio communications "space lab," and launched to GEO altitude. He called the modified LM the LM Relay Experiment Laboratory (LM REL).

A 1966 Lockheed proposal for a Lunar Module-derived radio experiment laboratory in the Apollo Applications Program.
Fordyce suggested that the LM REL's development and operation should occur within NASA's new Apollo Applications Program (AAP). Begun formally in August 1965, AAP aimed to apply spacecraft and technologies developed for Apollo lunar missions to new space missions of direct benefit to people on Earth.

The LM REL would be "visited periodically by crews to replenish, repair, install, initiate[,] and operate a variety of experiments," Fordyce wrote. Some of these experiments would "test the capability of a [GEO] relay to replace the aircraft, ships, and certain of the 30[-foot-diameter] antenna ground stations of the Manned Space Flight Network (MSFN)," he added.

Fordyce explained that, during Apollo missions, eight specially instrumented KC-135 aircraft, five tracking ships, and eleven 30-foot-diameter dish antennas would be needed to continuously link the Apollo spacecraft and the Mission Control Center in Houston, Texas. If a GEO communication satellite network replaced much of the MSFN, he wrote, then the result could be a "significant [cost] savings for NASA."

The "continuous contact capability" the GEO satellite network would provide would be especially valuable for missions in low-Earth orbit, Fordyce continued. Such missions spend only a few minutes above the horizon as viewed from any one Earth-surface antenna. Relaying transmissions from LEO spacecraft via GEO satellites would, he wrote, "permit greater flexibility in [spaceflight] operations by relaxing requirements to conduct difficult maneuvers [such as dockings and mission aborts] over instrumented sites."

Fordyce proposed two methods for placing the LM REL into its operational orbit (a Syncom 2-type synchronous orbit inclined 13.2° relative to Earth's equator). Both would see the LM REL launched with an Apollo Command and Service Module (CSM) spacecraft bearing a crew of three.

The most capable LM REL would rely on a three-stage Apollo Saturn V to reach GEO. The first two Saturn V stages would burn to depletion and fall away, then the S-IVB third stage would fire briefly to place itself, the CSM, and the LM REL into 100-nautical-mile-high low-Earth orbit (LEO). The S-IVB's J-2 engine would then be used to conduct three additional maneuvers over six hours to change the spacecraft's orbital inclination relative to the equator and increase its altitude.

Following the fourth S-IVB burn, the piloted CSM would separate, turn end for end, dock with a drogue docking unit on top of the LM REL, and withdraw it from the spent S-IVB stage. A CSM Service Propulsion System (SPS) main engine burn at GEO altitude would then nudge the LM REL into its operational orbit. The crew would enter the LM REL and conduct experiments for an unspecified period of time. After completing their mission, the astronauts would undock their CSM from the LM REL and ignite the SPS to return to Earth.

Alternately, the LM REL could climb from LEO to its GEO operational orbit on its own, though at the cost of reduced capabilities. Fordyce did not mention it — no doubt he expected that his audience would need no explanation — but the alternate approach would eliminate the need to develop an S-IVB J-2 engine capable of starting four times (the S-IVB J-2 for Apollo lunar missions was designed to ignite only twice).

The LM REL and a piloted CSM would reach 100-nautical-mile LEO together on a Saturn V or separately on a pair of two-stage Saturn IB rockets. In both cases, the CSM would dock with the LM REL and extract it from the S-IVB stage that boosted it into LEO.  The crew on board the CSM would then ready the LM REL for operations. Their work completed, the crew would undock in the CSM; the LM descent stage engine would then ignite to begin the LM REL's 5.25-hour climb to geosynchronous orbit.

As the LM REL reached GEO altitude, the spent descent stage would separate, then the LM REL ascent stage engine would ignite to complete insertion into operational orbit. Fordyce called the ascent-stage-only LM REL a "prototype" lab.

Artist's impression of first-generation TDRS satellite. Image credit: NASA.
Deep cuts in AAP's budget contributed to NASA's decision not to take up Fordyce's proposal. NASA did, however, eventually establish a GEO communications satellite network that replaced much of the MSFN, just as Fordyce suggested. Since the late 1980s, the Tracking and Data Relay Satellite System (TDRSS) has permitted near-continuous communications between controllers on the ground and the International Space Station, Hubble Space Telescope, and other LEO spacecraft.

The first satellite in the TDRSS network, the 2268-kilogram Tracking and Data Relay Satellite (TDRS)-1, reached LEO on 4 April 1983 on board the Shuttle Orbiter Challenger. A malfunctioning solid-propellant rocket stage failed to boost TDRS-1 all the way to GEO; controllers were able, however, to use the satellite's small attitude-control thrusters to nudge it into GEO over a period of about three months.

At launch, TDRS-1 was expected to operate for seven years, but it continued as part of the TDRSS network until October 2009, when its last remaining amplifier malfunctioned. In 2010, controllers used its attitude-control thrusters to boost it to a graveyard orbit about 483 kilometers (300 miles) above GEO altitude.

The second TDRS satellite was lost with the Orbiter Challenger and its seven-person crew during Space Shuttle mission STS 51-L (28 January 1986). The Space Shuttle launched five more first-generation TDRS satellites in 1988, 1989, 1991, 1993, and 1995. TDRS-3, the second to reach GEO, was repositioned to take over from TDRS-1. It is the oldest operational TDRS satellite. TDRS-4 was retired in December 2011 and placed in graveyard orbit in April 2012.

Three second-generation TDRS satellites, launched on Atlas IIA expendable rockets, reached GEO in 2000 and 2002. All remain operational.

NASA launched the third-generation TDRS-11, TDRS-12, and TDRS-13 satellites on Atlas V 401 rockets in 2013, 2014, and 2017, respectively. They also remain operational.

Launch of the 3454-kilogram (7615-pound) TDRS-13 satellite on 18 August 2017. Image credit: NASA.


Memorandum with attachment, MLO/Samuel Fordyce, SAA Flight Operations, to MLD/Deputy Director, Saturn/Apollo Applications and MLA/Director, Apollo Applications, AAP Synchronous Mission, April 29, 1966.

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