Engineer Special Study of the Surface of the Moon (1960-1961)

Engineer Special Study Sheet 1: Generalized Photogeologic Map of the Moon. Please click to enlarge. Image credit: USGS.
The race to the Moon began on 17 August 1958 and the Soviet Union won. This isn't the opening line of an alternate history story; rather, it is an acknowledgment that more than one Moon race took place. The first, with the goal of launching a small automated spacecraft to the Moon, began with the liftoff of the Able 1 lunar orbiter, a 38-kilogram U.S. Air Force (USAF) probe. (It was later re-designated Pioneer 0.) Able 1's first stage, a Thor missile, exploded just 77 seconds after launch from Cape Canaveral, Florida, ending the world's first attempted lunar mission.

A month later, on 23 September 1958, the Soviet Union joined the race. A spherical Luna probe intended to impact the Moon fell victim to the failure of its upgraded R-7 booster rocket just 93 seconds after liftoff from Baikonur Cosmodrome in central Asia.

On 11 October 1958, the USAF launched Able 2, a near-copy of Able 1. It was the first lunar launch conducted under NASA auspices. The civilian space agency had opened its doors on 1 October 1958. NASA absorbed most Department of Defense space projects, though in practice the USAF and U.S. Army continued to carry out missions while interagency relations and lines of command became defined.

Able 2, later re-designated Pioneer 1, burned up in Earth's atmosphere on 13 October after its Able rocket second stage shut down early, placing it on an elliptical path that took it about a third of the way to the Moon. The Soviets launched their second Luna Moon impactor just 16 hours after the U.S. launched Able 2. The unnumbered Luna's upgraded R-7 launch vehicle exploded 104 seconds after liftoff.

And so it went, with launches from Florida and Kazakhstan alternating and failing. The Pioneer 2 lunar orbiter (8 November 1958) and another Luna impactor (4 December 1958) fell victim to premature launch vehicle shutdowns. Pioneer 3 (6-7 December 1958), the first NASA/Army Moon probe, was launched on a U.S. Army Juno II, not a USAF Thor-Able, but performed much as had Pioneer 1.

First attempt: Thor-Able 1 launches Pioneer 0 (17 August 1958). Image credit: Air Force Air & Space Museum. 
On 3 January 1959, the Soviet Union snatched victory from the jaws of defeat. Their Luna 1 impactor missed the Moon by 6400 kilometers, and so failed to accomplish its mission. It sailed on, however, becoming the first human-made object to orbit the Sun. The Soviets nicknamed it Mechta ("dream"). The U.S. Army launched the Pioneer 4 lunar flyby spacecraft two months later (3 March 1959). It failed to return images of the Moon, but repeated Mechta's feat.

Another unnumbered Luna impactor fell victim to an R-7 failure on 18 June 18 1959. Then, on 14 September 1959, on their sixth attempt, Soviet rocketeers succeeded in striking the Moon with the Luna 2 impactor. The probe struck near the center of the Moon's Nearside, the hemisphere that faces the Earth. Three weeks later (6 October 1959), Luna 3 flew 7900 kilometers over the Moon's south pole and imaged the hidden Farside hemisphere.

In a last-ditch effort to steal the Soviet Union's thunder, the USAF and NASA decided to give a planned Pioneer Venus orbiter a new mission: orbit and photograph the Moon at close range. Its mission ended 104 seconds after liftoff on 26 November 1959, when its Atlas-Able launcher lost its streamlined launch shroud and tumbled out of control.

As the first Moon race ended in Soviet victory, pressure built in the U.S. for a rematch. Though President Dwight Eisenhower had made it clear that the Department of Defense branch services should concentrate on space and rocket projects with immediate military applications, the Moon still beckoned to U.S. Army and USAF rocketeers.

The U.S. Army and the USAF studied lunar surface bases even after the creation of NASA. The Army Ballistic Missile Agency emphasized Project Horizon, a lunar fort, while the USAF worked with contractors on the SR-183 Lunar Observatory project. LUNEX was a USAF study of an early manned lunar expedition. The USAF also began lunar mapping using Earth-based telescopes.

Moon fort: Project Horizon lunar base. In this painting from 1959, a U.S. Army crew lander arrives at the landing field in the background, beyond which lies a jagged line of mountains. In the foreground, habitat modules are buried in an excavated ditch for micrometeoroid protection. Image credit: National Air & Space Museum.
The first attempt to map lunar features for scientific and engineering purposes did not, however, originate within the Defense Department. It was begun instead by Arnold Mason of the U.S. Geological Survey (USGS) Military Geology Branch in Washington, DC. According to Don Wilhelms, writing in his 1993 memoir To a Rocky Moon, the peripatetic Mason became interested in lunar geology after the 4 October 1957 launch of Sputnik 1. Mason's boss, Frank Whitmore, soon got caught up in his enthusiasm. Whitmore, incidentally, served as Secretary of the Geological Society of Washington.

Early in 1959 — soon after Luna 1 — Mason proposed to carry out an analysis of the Moon's alien terrains to determine their suitability for spacecraft landings, travel on foot and by rover, and base construction. With Whitmore's blessing, he enlisted Robert Hackman and Annabel Brown Olson of the USGS Photogeology Branch in his project. Mason became project chief, Hackman became Mason's co-author, and Olson (who, according to Wilhelms, received insufficient credit for her labors) assisted Hackman. At first, they had available only meager USGS funds. Soon after Luna 2 and Luna 3, however, the Army Corps of Engineers funded their study.

Mason and Hackman's assessment took in only the Nearside. They based their analysis on photographic plates from large telescopes on Earth, which under the best viewing conditions could (they estimated) reveal features on the moon no smaller than about a mile across. In fact, features 10 miles wide were barely discernible in most of the photographic images they used.

Their work soon drew in as consultants lunar experts Gerard Kuiper (McDonald Observatory), Eugene Shoemaker (USGS Menlo Park), and Robert Dietz (Naval Electronics Laboratory). All three supported the impact hypothesis, which stated that most of the Moon's craters are asteroid impact scars; not, as some believed, volcanic calderas. At the time, planetary astronomer Kuiper was hard at work on a USAF-funded lunar photographic atlas; Mason and Hackman would use it near the end of their study. Shoemaker, meanwhile, was busy refining a prototype lunar geologic map of the region containing the large, relatively young crater Copernicus; Hackman would later assist him with identification of lineaments in the Copernicus region.

The Army Corps of Engineers published the first edition of Mason and Hackman's four-sheet "Engineer Special Study of the Surface of the Moon" map set in July 1960. USGS published a second edition with "minor revisions" the following year.

The "Engineer Special Study" was significant in part because its Sheet 1 (top of post), titled "Generalized Photogeologic Map,” was the first major lunar map to show stratigraphic relationships: that is, it attempted to display the chronological order of the formation of the Moon's surface features. Mason and Hackman's stratigraphic system centered on the formation of the maria (Latin for "seas"), the relatively smooth, dark-hued plains that mottle the Nearside. They make up about 20% of the Moon's surface.

Mason and Hackman colored orange the heavily cratered, light-colored "pre-maria" terrain; that is, landforms that they believed were already in place when the maria formed. They colored maria yellow, while green indicated "post-maria" features; mainly young asteroid impact craters, but also features that they interpreted as being of recent volcanic origin. They used black dots to mark what they identified as volcanic cones and domes and thin black lines to mark what they thought were tectonic faults.

Their stratigraphic map, though pioneering, was too simplistic to accurately portray the Moon's history. Most of the maria basins formed at different times during the first billion or so years of lunar history, so features associated with them often overlap. An impact crater blasted into an older mare (Latin for "sea") would, for example, become a pre-maria landform by Mason and Hackman's reckoning if it became engulfed in ejecta and lava from a later basin-forming giant impact. In addition, some prominent lunar features identified as pre-maria (the Apennine Mountains, for example) should have been represented by a fourth color to signify that they are non-maria features created by the same giant asteroid impacts that excavated the maria basins.

By contrast, Shoemaker's nearly contemporaneous prototype Copernicus geology map, printed in small quantity by the USAF Aeronautical Chart and Information Center in April 1961 but never formally published, identified five stratigraphic "systems." From oldest to youngest, these were the Pre-Imbrian, Imbrian, Procellarian, Erastothenian, and Copernican systems. Even this would turn out to be simplistic, however, once robot and human explorers began to provide lunar geologists with close-up images and samples of the Moon's complex terrain.

Engineer Special Study Sheet 2: Lunar Rays. Please click to enlarge. Image credit: USGS.

Engineer Special Study Sheet 3: Physiographic Divisions of the Moon. Please click to enlarge. Image credit: USGS.
In sheet 2 of the "Engineer Special Study," titled "Lunar Rays," Mason and Hackman plotted the source craters and extent of the Moon's most prominent ray systems. They correctly identified the light-hued rays as ejecta blasted out from young asteroid impact craters.

Mason and Hackman's Sheet 3, titled "Physiographic Divisions of the Moon," was their most ambitious. In it, they applied photogeologic principles pioneered on Earth to identify more than 70 different lunar terrain units.

Sheets 1 through 3 laid the groundwork (literally) for Sheet 4, on which Mason assessed in writing the landing, travel, and construction conditions in each of the physiographic regions on Sheet 3. What follows are summaries of his assessments for several regions that have been visited by spacecraft.

Luna 2 struck the southern flank of Autolycus crater in the northern part of Mason and Hackman's Apennines Region. According to Mason and Hackman's analysis, Autolycus is a post-maria impact crater, only lightly rayed, on the western edge of Mare Imbrium, in the extensive Mid Lunar Lowlands. Mason wrote that the surface in the Apennines Region is rough and blocky, so landings there would be very difficult. Movement in the region would, he judged, be the "most difficult on the [M]oon's surface, and possible only by carefully selected routes." Construction would be "very difficult because of blocky material and steep slopes."

James Irwin salutes Old Glory at Hadley-Apennine in a photograph captured by Apollo 15 Commander David Scott. The Lunar Module Falcon and the Lunar Roving Vehicle Scott and Irwin used to explore the Hadley-Apennine site glitter in the harsh morning sunlight. The surface material around Falcon is rolling and loose with few large rocks. Mount Hadley Delta, about 4000 meters tall and rounded by billions of years of small meteoroid impacts, stands behind Irwin and Falcon. Image credit: NASA.
Luna 2 was not designed to return images as it plunged toward the Moon; however, the Apollo 15 Lunar Module Falcon landed west of the Luna 2 impact site on July 30, 1971. Astronauts David Scott and James Irwin found the area to be cratered and rolling, but difficult neither to land on nor to navigate on foot or by rover. The surface material was loose to a depth of many meters. The nearby Apennine Mountains, which Mason and Hackman had envisioned as steep and jagged, turned out to have been rounded and partly leveled by micrometeoroid impacts over the nearly four billion years since their formation.

NASA's Ranger 7 probe was designed to return images of the lunar surface as it fell toward destructive impact. On 31 July 1964, Ranger 7 returned more than 4300 photos of the area between Oceanus Procellarum and Mare Nubium. 

Mason and Hackman had called the area containing Ranger 7's impact site the Riphaeus Section. It was a lowland maria divided by the highland Riphaeus Mountains. Mason judged that landing and movement would be "generally easy" if blocky isolated pre-maria highland areas and post-maria craters could be avoided.

Mind the crease: the Riphaeus Section from Sheet 3. Ranger 7 impacted the Moon southwest of the heavily degraded Fra Mauro crater, which is marked by a dashed outline at center right. Please click to enlarge. Image credit: USGS.
Construction, on the other hand, would be a challenge in the Riphaeus Section. Mason expected that, under a thin layer of loose debris, lunar base builders would find basaltic rock hard enough to prevent boring and excavation. Whereas in the Apennines Region he advised lunar base builders to avoid craters and their blocky surroundings, in the Riphaeus Section such asteroid-shattered areas would probably be the only places where digging could occur. This applied to other maria lowlands as well. 

Scientists examining Ranger 7 images found that its impact area was cratered down to the scale of inches; however, the craters were almost all eroded, with smooth floors and rims and few large rocks. Micrometeoroids had been whittling away at the terrain in the Riphaeus Section for a very long time. In tribute to Ranger 7, lunar mappers named the area where it impacted Mare Cognitum, which means "Known Sea."

Surveyor 7, the last of its series of soft landers, alighted gently on the northern flank of Tycho crater on 10 January 1968. Mason and Hackman identified the area containing post-maria Tycho as the pre-maria Macrocrater Province. Tycho, they wrote, spanned 54 miles from rim to rim. The crater's floor was 12,000 feet below its rim, which stood 7900 feet above the surrounding terrain. They noted that Tycho was the Moon's most prominent ray crater, with bright streaks extending up to 500 miles plainly visible to the unaided eye at full moon.

Closeup of Sheet 2: Tycho and the adjoining Macrocrater Province. Please click to enlarge. Image credit: USGS.
Mason judged that landing and movement would be difficult near Tycho. The latter would be possible, however, if a safe travel route could be selected in advance. Construction would be difficult because of the many large blocks embedded throughout the area.

Surveyor 7 landed blind on Tycho's flank; that is, it included no hazard-avoidance system. Through its scanning camera scientists saw that the area was indeed rougher than those that previous Surveyors had explored. They saw loose rocks, boulders, relatively steep slopes, apparent bedrock outcrops, and odd "lakes" of dark gray material, possibly cinders laid down by recent volcanism or rock melted by the colossal energies of the Tycho impact. Some of these features could have destroyed Surveyor 7 had it landed on them.

In general, however, Tycho, like the Riphaeus Section and the Apennines Region, was not as rugged as Mason had predicted. In fact, after Surveyor 7, some felt that Tycho's flank was smooth and level enough for Apollo astronauts to visit. A 1969 study based on Surveyor 7 images determined that it was too rough for rover operations, however.

Tycho's rocky flank: the view from Surveyor 7. Image credit: NASA.
In early December 1960, Mason and Hackman attended the International Astronomical Union's First Lunar Symposium at Pulkovo Observatory in Leningrad. The meeting was held in the Soviet Union in deference to that country's demonstrated lead in lunar exploration. They displayed the U.S. Army Corps of Engineers edition of the "Engineer Special Study."

Upon his return from the historic symposium, Mason presented an informal report on the trip to the January 1961 meeting of the Geological Society of Washington. Mason's boss Whitmore briefly summarized his report in the meeting minutes.

Hackman appeared as co-author on Shoemaker's April 1961 prototype Copernicus geologic map. Copernicus mapping then stalled for several years because Shoemaker had new responsibilities. He had succeeded in launching the NASA-supported Astrogeology Studies Project at USGS Menlo Park, near San Francisco, in August 1960; this became the NASA-supported USGS Branch of Astrogeology in September 1961. In addition, he was busy publishing ground-breaking papers on lunar cratering dynamics and lunar and terrestrial geologic timescales. The Copernicus map was eventually published in 1967 with soon-to-be-astronaut Harrison Schmitt and Newell Trask as Shoemaker's co-authors.

In July 1961, Hackman submitted for review what became after the "Engineer Special Study" the second published USGS lunar map: a geologic study of the Kepler region based on Shoemaker's lunar geologic mapping conventions and five-system lunar stratigraphic column. The Kepler map, published in 1962 under the auspices of the Branch of Astrogeology, was the first NASA-funded USGS lunar map to be published.

Eleven months after the Pulkovo symposium, in November 1961, Whitmore had the sad duty of informing the Geological Society of Washington of Mason's untimely death. The pioneering lunar mapper had taken his own life on 31 October 1961. He was 54 years old.

In his memoir, Wilhelms wrote that Mason committed suicide "for reasons that are not entirely clear and are undoubtedly complex, but which seem to have included non-recognition for his original and ardent pioneering of lunar studies for the U.S. Geological Survey." Pulkovo had marked the high point of Mason's lunar career: after that, Shoemaker's new program increasingly sidelined USGS lunar studies in Washington, DC.

Hackman's involvement in lunar geologic mapping was by then also drawing to a close. His steadfast refusal to leave the Washington area proved to be career limiting. Shoemaker transplanted the Branch of Astrogeology from Menlo Park to the small town of Flagstaff, Arizona, during 1963, and soon the name "Flagstaff" became synonymous with lunar and planetary mapping. Hackman completed one more map for the Branch of Astrogeology — a geologic map of the Moon's Apennines region, which was published in 1966 — but his pioneering contributions to lunar geologic mapping ceased with publication of the Kepler map.

Although the "Engineer Special Study" remained relatively obscure — and became even more so after data from lunar spacecraft rendered much of it obsolete — it did manage to earn a small place in popular culture. Chapter 12 of Arthur C. Clarke's 1968 novel 2001: A Space Odyssey, titled "Journey by Earthlight," begins with a description of the Macrocrater Province and the crater Tycho extracted from Mason's Sheet 4 of the "Engineer Special Study."

References

"Engineer Special Study of the Surface of the Moon," Robert J. Hackman and Arnold C. Mason, Army Map Service, Corps of Engineers, July 1960.

"Engineer Special Study of the Surface of the Moon," Miscellaneous Geologic Investigations Map I-351, Robert J. Hackman and Arnold C. Mason, U.S. Geological Survey, Washington, DC, 1961.

"Memorial to Arnold Caverly Mason (1906-1961)," H. Foster, Geological Society of America Bulletin, Vol. 73, August 1962, pp. 87-90.

To A Rocky Moon: A Geologist's History of Lunar Exploration, Don E. Wilhelms, The University of Arizona Press, 1993, pp. 37-42.

More Information

Around the Moon in 80 Hours (1958)

"Essential Data": A 1963 Pitch to Expand NASA's Robotic Exploration Programs

Apollo Science and Sites: The Sonett Report (1963)

An Apollo Landing Near the Great Ray Crater Tycho (1969)

Log of a Moon Expedition (1969)

Could the Voyages in the Film and Novel "2001: A Space Odyssey" Really Happen? (Part 1)

Integral Launch and Reentry Vehicle: Triamese (1968-1969)

Triamese target: a large Earth-orbital "Space Base" assembled from modules launched atop two-stage Saturn V rockets. The Space Base, expected to be operational by about 1980, would be staffed by up to 100 people. Image credit: NASA.
The Triamese concept originated in 1967 in a reusable launch and reentry vehicle study General Dynamics Convair (GDC) performed on contract to the U.S. Air Force (USAF). Triamese owed its peculiar name to its peculiar launch configuration. At liftoff it would comprise one orbiter element and two booster elements. The boosters would together serve as the first stage; they would also provide propellants to the orbiter's engines during first-stage boost. One booster would attach to the orbiter's flat belly and the other to its rounded back. 

Space launch vehicle concepts with separate reusable booster and orbiter elements were not exactly new in 1967. What was different about Triamese was its strict reliance on a common booster and orbiter design. The Triamese orbiter and booster elements were intended to be virtually identical. GDC explained that

[i]n order to achieve the economy predicted for the Triamese system, the orbital and boost elements must have a high degree of commonality and must represent essentially a single development program. . .This commonality has been obtained by "overdesigning" the boost elements. . .[which] creates performance penalties that are accepted.

GDC called Triamese "a new mixture of aircraft, spacecraft, and launch vehicle." The Initial Point Design (IPD) Triamese launch stack (A, above) would have comprised two booster elements and one orbiter element, all virtually identical. It would have measured 149.5 feet (45.6 meters) tall from the trailing tips of its six rudder fins (two per element) to its three noses. B, a tail-on view of one element, shows the V-shaped, 46.1-foot (14-meter) spread of the rudder fins, 21-foot-wide (6.4-meter-wide) flat belly, and twin XLR-129 rocket engines arranged one above the other. Turning view B 45° horizontally yields view C. The IPD Triamese element would measure 31.4 feet (9.6 meters) from its belly to the tops of its rudder fins. View D displays "switchblade" wings deployed for stable subsonic flight. Wingspan is 107.5 feet (32.8 meters). Image credit: General Dynamics Convair/DSFPortree
The Triamese concept helped to shape NASA's May 1968 Integral Launch and Reentry Vehicle (ILRV) study Statement of Work and the ILRV Request for Proposal the space agency released to U.S. industry in October 1968. When time came for NASA to select four industry proposals for ILRV study contracts in January 1969, it was a foregone conclusion that Triamese would be counted among them.

NASA Marshall Space Flight Center (MSFC) in Huntsville, Alabama, was tasked with managing the GDC ILRV study contract. NASA MSFC was home of the three-stage Apollo Saturn V rocket. At the time of the ILRV study, Apollo Saturn V development, manufacture, and testing were drawing to a close. Managers at the Huntsville center hoped, however, that a two-stage Saturn V variant designated INT-21 might launch a series of increasingly complex space stations in the 1970s.

INT-21 consisted of the first two stages of the Saturn V — the S-IC first stage and S-II second stage — both of which measured 33 feet (10 meters) in diameter. An Earth-orbital payload measuring up to that diameter — for example, a large space station module — would replace the 21.7-foot-diameter (6.6-meter-diameter) S-IVB third stage of the Apollo Saturn V. 

One station program scenario, favored by NASA Administrator Thomas Paine, would see INT-21-launched Apollo Applications Program (AAP) Orbital Workshops — converted S-IVB stages — lead in 1975 to a large drum-shaped station with up to 12 crewmembers. Multiple INT-21-launched large station modules might then be joined together in orbit as early as 1980 to form a "Space Base" with up to 100 staff.

In that scenario, the ILRV shuttle would serve as a Saturn V supplement. The big rocket would do the heavy lifting all through the 1970s, leaving to the smaller reusable shuttle the specialized task of affordably launching astronauts, supplies, replacement parts, and scientific experiment apparatus to the space station and returning astronauts, experiment results, and data products to Earth. 

GDC began its ILRV Triamese study with an Initial Point Design (IPD) based on its USAF study results and inputs from NASA engineers. The IPD Triamese was designed to deliver up to 25,000 pounds (11,340 kilograms) of supplies and equipment to the space station and return up to 2500 pounds (1130 kilograms) to Earth during a single flight. The two boosters and the orbiter would each carry a flight crew of two astronauts, for a total of six. In addition, the orbiter would include a passenger compartment for transporting 10 astronauts to and from the space station. 

Orbiter and booster commonality was not the only cost-saving principle underpinning the IPD Triamese system. Another was use of off-the-shelf technology. GDC proposed, for example, that the design of the Triamese "switchblade" wings, which would enable stable flight at subsonic speeds, should be based on the variable-geometry wing system of the F-111 "Aardvark" aircraft the company manufactured for the USAF. 

The variable-geometry wings of the supersonic F-111 in action. In 1967, the F-111 became the first variable-geometry aircraft to enter active service. Image credit: U.S. Air Force.
GDC envisioned that the IPD Triamese elements would, like operational airplanes, fly repeatedly with minimal refurbishment between flights. The company acknowledged, however, that the elements would be subjected to greater stress during flight than would most aircraft, leading to greater potential for component failure.

GDC proposed to solve this problem by equipping IPD Triamese subsystems with sensors linked to on-board magnetic-tape flight recorders. After landing, data on subsystem performance would be carefully analyzed. Hardware that showed signs of actual or impending trouble would be subjected to detailed inspection and possible repair or replacement. 

The sensors would also enable a detailed on-board checkout capability that would slash costs by allowing NASA to get by with only a simple launch control center. KSC's Apollo Saturn launch control center was expansive and expensive, with many control consoles and an army of highly trained personnel; IPD Triamese launch control might more closely resemble an airport control tower. 

GDC expected that the IPD Triamese design, development, and test program would begin on 1 November 1971 and last until the first operational IPD Triamese flight on 1 January 1977, a period of 62 months. Engineering design would occur between 1 November 1971 and 1 July 1974. Development of the Pratt & Whitney XLR-129 rocket engine, which GDC called a "pacing item," would last from 1 November 1971 to 1 August 1974. Rocket engine tests using IPD Triamese vehicles that were captive  — that is, bolted down so that they could not take off — would take place between 1 March 1975 and 1 March 1976.

GDC proposed a "fatigue test vehicle" to help to ensure that the IPD Triamese elements would be as reusable as expected. This would take the form of a skeletal IPD Triamese element with all systems installed except for the metal plates and insulation blankets of its heat shield. 

Beginning on 1 November 1974, the fatigue test vehicle would undergo repeated propellant tank and cabin pressurizations, switchblade wing, turbofan jet engine, and landing gear deployments, computer starts, and other subsystem activations so that engineers could gain insight into malfunction characteristics and operational lifetimes. The tests would continue into the period of operational IPD Triamese flights.

The Initial Point Design (IPD) Triamese orbiter element differed from the booster element only in detail. Unless otherwise noted, all features called out in the side view drawing above are features of both the orbiter and the booster. A: cockpit for two astronauts seated side by side; B: passenger compartment for 10 Space Station crewmembers with seating arranged in three rows (orbiter only); C: forward landing gear (stowed). D: forward landing gear (down and locked); E: short liquid oxygen tank (orbiter); F: leeward forward pin connection (orbiter only); G: windward forward pin connection; H: cargo bay hatch (orbiter only); I: cargo bay (orbiter only); J: main landing gear (stowed); K: main landing gear (down and locked); L: short liquid hydrogen tank (orbiter only); M: switchblade wing compartment; N: leeward propellant feeds; O: windward propellant feeds; P: XLR-129 rocket engine (one of a pair); Q: body flap with elevons; R: rudder fin (one of a pair); S: rudder flap (one of a pair); T: extendible engine skirt (orbiter only — retracted); U: extendible engine skirt (orbiter only — extended). Image credit: General Dynamics Convair/DSFPortree.

Top view of IPD Triamese element. Unless otherwise noted, all features called out in the top view drawing above are features of both the orbiter and the booster. 1: cockpit windows; 2: cockpit crew hatch; 3: passenger compartment crew hatch/docking unit (orbiter only); 4: turbofan jet engine (stowed); 5: turbofan jet engine (deployed and locked); 6: long liquid oxygen tank (booster only); 7: reinforcing ring for attachment of forward pin connection (booster only) or connections (orbiter only), landing gear, and switchblade wing pivot; 8: switchblade wing pivot (one of a pair); 9: switchblade wing (deployed — one of a pair); 10: switchblade wing flap (one of a pair); 11: switchblade wing (stowed — one of a pair); 12: main landing gear (stowed); 13: cargo bay hatch/docking unit (orbiter only); 14: long liquid hydrogen tank (booster only); 15: aft attachment pin actuator (booster only); 16: leeward propellant feeds (one of a pair); 17: rudder fin (one of a pair); 18: rudder flap (one of a pair); 19: non-extendible XLR-129 rocket engine skirt (booster only); 20: body flap with elevons. Image credit: General Dynamics Convair/DSFPortree.

IPD Triamese flight testing would use "an aircraft approach." All flights would carry two test pilots per element — there would be no unpiloted IPD Triamese test flights. GDC allotted three booster elements and three orbiter elements for the IPD Triamese test program. Of these, two boosters and one orbiter would be carried over to operational flights. 

GDC scheduled 50 horizontal test flights at Edwards Air Force Base, California, between 1 October 1974 and 1 March 1976. During these tests, individual IPD Triamese elements would use their twin TF-34 turbofan jet engines to take off from a runway with their switchblade wings extended to verify subsonic flight and landing characteristics. 

The General Electric-built TF-34 engine generated 12,600 pounds (5715 kilograms) of thrust. GDC was familiar with the engine because it used it in its proposal for the U.S. Navy's S-3 Viking aircraft. The engine produced a characteristic low rumble, a sound that would no doubt have become associated with piloted spaceflight had NASA given GDC the nod to build the IPD Triamese.

A U.S. Navy S-3 Viking aircraft descends to a carrier landing. Visible is one of its two General Electric-built TF-34 jet engines. The IPD Triamese shuttle orbiter and booster elements would each have included two such engines. In the unlikely event that a returning IPD Triamese element missed its first attempt at a landing on the runway at NASA Kennedy Space Center, the jet engines would have permitted a second try. Image credit: U.S. Navy.
The company scheduled 15 single-element rocket-propelled vertical flights at NASA Kennedy Space Center (KSC) on Florida's east coast between 1 September 1975 and 1 November 1976. The tests would, among other things, enable verification of IPD Triamese flight characteristics at transonic and supersonic speeds. 

The IPD Triamese element under test would lift off from one of two launch pads built at KSC specifically for IPD Triamese launches, climb to a specified altitude, and shut down its twin rocket engines. It would then pitch over to horizontal attitude, deploy its wings and jet engines, and fly to a runway at KSC built specifically for IPD Triamese landings. 

In December 1975, the flight test program would shift into high gear as preparations began for suborbital two-element test flights, the first IPD Triamese flights to launch astronauts into space. A pair of joined booster elements would lift off vertically from a KSC IPD Triamese pad on 15 February 1976, separate, and undergo a reentry virtually identical to that they would experience during operational Triamese flights. They would then land on the KSC IPD Triamese runway. NASA would repeat this test on 1 April 1976. 

About two weeks later, on 15 April 1976, the first booster-orbiter suborbital flight test would take place. It would closely resemble the booster-booster tests. The second booster-orbiter test would occur on 1 June 1976. 

The IPD Triamese flight test series would end with a pair of three-element orbital flight tests on 1 August and 1 November 1976. The missions would see the first IPD Triamese dockings with an Earth-orbiting space station. 

The boosters and orbiter flown during the second orbital test flight would be used for "refurbishment verification" — a rehearsal of the normal IPD Triamese post-flight checkout and maintenance "turnaround" process — then the orbiter and one booster would be held in reserve as "standby elements" for the first operational flight of the IPD Triamese program on 1 January 1977.

Availability of standby elements — a backup orbiter and a backup booster — would be a standard part of preparation for every operational IPD Triamese mission. If an active orbiter or active booster suffered damage or malfunctioned and required time-consuming repairs, a standby element would fill in for it so that launch could go ahead as scheduled. This approach recognized the critical role reliable space transportation would play in NASA's space station program. 

GDC proposed that, in addition to the two standby elements, NASA's IPD Triamese fleet should include four active orbiters and six active boosters. The orbiters would each fly once per month, for a total of 48 orbiter flights per year. The boosters would each fly 16 times per year, for a total of 96 booster flights. 

Diagram of IPD Triamese orbiter and booster turnaround flow. In one month, four active orbiters would lift off from Kennedy Space Center, Florida. In the same period, four active boosters would fly once and two would fly twice. A fifth orbiter and a seventh booster would serve as "standby elements" ready to enter the turnaround flow if an active orbiter or booster should be grounded for repairs. Image credit: General Dynamics Convair/DSFPortree.

At the start of every operational IPD Triamese mission, turnaround technicians would load the 17.5-foot-diameter (5.3-meter-diameter), 12.4-foot-long (3.8-meter-long) payload bay located between the orbiter's liquid oxygen tank and its liquid hydrogen tank with 25,000 pounds (11,340 kilograms) of supplies and equipment bound for the Space Station. The orbiter propellant tanks would be made shorter than the booster tanks to make room for the 3000-cubic-foot (85-cubic-meter) bay.

Turnaround technicians would next pump consumables into the IPD Triamese elements. These would include 4660 pounds (2110 kilograms) of jet fuel for each booster and 1610 pounds (730 kilograms) for the orbiter, along with 3820 pounds (1730 kilograms) of attitude control propellants for the orbiter and 1420 pounds (645 kilograms) for each booster. 

The three elements would then be towed to the launch pad on their extended tricycle landing gear, hoisted vertical, and, after their landing gear was retracted, mounted on the pad on three support struts each. After the vehicles were joined to each other by three "pin connections," one forward and two aft, five support struts (the three supporting the orbiter and one each supporting the boosters) would be removed, leaving in place two per booster. 

Launch pad technicians would connect propellant feed lines linking the orbiter and the booster propulsion systems and attach umbilical hoses for propellant tank loading. After a leak check using on-board checkout equipment, they would fill the orbiter's tanks with 362,800 pounds (164,560 kilograms) of liquid oxygen and 51,830 pounds (23,510 kilograms) of liquid hydrogen. Each booster would be loaded with 424,500 pounds (192,550 kilograms) of liquid oxygen and 62,890 pounds (28,525 kilograms) of liquid hydrogen. Before vacating the vehicles, the pad technicians would conduct a final check of the propulsion system using on-board checkout equipment. 

The three flight crews and passengers would board, then the flight crews would perform a final check of all on-board systems save propulsion. Finally, at a time selected to enable a quick rendezvous with the Space Station, the six XLR-129 engines would ignite and power up to 20% of maximum sea-level thrust. There they would briefly hold to allow the flight crews to check engine performance. If all six engines were found to be operating normally, they would power up to 100%, hold-down attachments on the four support struts would disconnect, and the IPD Triamese stack would lift off.

IPD Triamese launch and ascent: the IPD Triamese launch stack (A) would stage at an altitude of 160,000 feet (48,770 meters) (B). The twin boosters would undergo a low-stress suborbital reentry (C), then would level off at 15,000 feet (4570 meters). Their flight crews would extend their jet engines and wings, then fly back in tandem to their NASA KSC base (D), a distance of 185 nautical miles (340 kilometers). The orbiter, meanwhile, would continue its journey (E) to the Space Station in 270 nautical-mile (500-kilometer) low-Earth orbit. Image credit: General Dynamics Convair/DSFPortree.

At liftoff, the four booster engines would each generate 394,500 pounds (178,715 kilograms) of thrust; the two orbiter engines, 380,000 pounds (172,365 kilograms) each. GDC calculated that the IPD Triamese stack would weigh 1,751,000 pounds (794,240 kilograms) at liftoff. Of this, the boosters would each account for 596,450 pounds (270,545 kilograms) and the orbiter, 558,100 pounds (253,150 kilograms).

During the first stage of ascent, the twin booster elements would supply all propellants to their own engines and the two orbiter engines. GDC did not specify how long first-stage flight would last. The company calculated, however, that the entire journey from launch pad to orbit would last only 6.2 minutes. Acceleration during ascent would top out at four times the pull of Earth's gravity.

GDC assumed that NASA's space station destination would circle the Earth in an orbit inclined 55° relative to Earth's equator. IPD Triamese launch azimuth would, however, be set at 35° to avoid overflight of the U.S. east coast early in the ascent phase. This meant that the orbiter would have to perform a westward yaw ("dogleg") maneuver to reach 55° orbit.

GDC estimated that flight conditions during ascent were 500 times more likely to cause a system failure than were conditions in space. As might be expected, engines, propellant feeds, and avionics were the systems most likely to malfunction. The company cited possible failure modes virtually certain to lead to structural failure and loss of life in as little as one second — for example, a hydraulic system failure that would cause the engines of one of the three elements to gimbal (pivot) and lock suddenly. 

To avoid such catastrophic failures, GDC proposed automatic malfunction detection and switchover to backup systems. This approach would, the company estimated, reduce the IPD Triamese catastrophic failure rate to one in 2000 flights.

Switching to backups might allow an IPD Triamese mission to proceed as normal. Even if an abort were necessary, under most circumstances the boosters would return to the KSC runway as normal. The orbiter, on the other hand, might seek to return directly to KSC, reach a low orbit and return to KSC after circling the Earth once (the generally preferred option), bank eastward and land downrange on the North Atlantic island of Bermuda, or, in the worst-case scenario, ditch at sea or crash-land on the Arctic ice cap. 

Booster thrust per engine would increase to 433,300 pounds (196,540 kilograms) just before burnout. The orbiter engines, meanwhile, would each extend an expendable skirt just before staging, allowing an increase in thrust per engine to 460,500 pounds (208,880 kilograms). 

The boosters would expend their propellants as the IPD Triamese stack reached a speed of 6800 feet per second (2070 meters per second). After booster separation, thrust per orbiter engine would steadily decrease until it reached 310,000 pounds (140,620 kilograms) just before shutdown. 

After they separated from the orbiter, the boosters would perform a suborbital reentry and turn toward KSC. They would deploy their switchblade wings and jet engines and fly back to base at a speed of 225 miles (365 kilometers) per hour. 

Staging during ascent to orbit: the operations illustrated above would last no longer than nine seconds. The orbiter (A) is shown with twin XLR-129 engines firing and engine skirts extended. Pyrotechnic bolts would fire in the booster (B) forward pin connections, allowing aerodynamic drag and inertia to cause the boosters to tip away from the orbiter. C: aft pin connection actuators on the boosters simultaneously extend to ensure adequate clearance between the booster body flaps and the orbiter engine bells. D: when the boosters tipped back to an angle of 20° relative to the orbiter center line, pyrotechnic bolts sever the two aft pin connections. E: the aft pin connection actuators on the boosters retract. The boosters would then roll to turn their windward sides toward their direction of flight and begin descent and return to NASA Kennedy Space Center. Image credit: General Dynamics Convair/DSFPortree.

GDC proposed an IPD Triamese Reaction Control System (RCS) with 24 nitrogen tetroxide/hydrazine thrusters, most of which would cluster near the nose and tail. Of the 24, half would generate 1420 pounds (644 kilograms) of thrust and half 1160 pounds. 

Eight of the former would serve as orbital maneuvering thrusters, with four facing forward and four aft. These would permit the orbiter flight crew to circularize their orbit at space station altitude and perform rendezvous and station-keeping with the station. The company noted that the eight orbital maneuvering thrusters could be omitted from the boosters if doing so would save money.

The IPD Triamese orbiter mission would last 25 hours. Of this, the orbiter would spend 17.3 hours attached to the space station, during which time it would rely on station electricity, attitude control, life support, and communications. 

Precisely how the orbiter would link up with the space station was not explained. The liquid oxygen tank would be located between the cargo bay and the passenger compartment, preventing movement between them; for this reason, each would require an exterior hatch. This implies the existence of two docking units, one for each hatch, or a station hangar surrounding both hatches that could be pressurized. Though drawings show the cargo bay hatch as round, GDC described it as square and five feet (1.7 meters) wide. 

The company also did not describe the method of cargo transfer. No doubt the transfer of 25,000 pounds (11,340 kilograms) of supplies and equipment to the space station would need to be carefully orchestrated if it was to be completed in 17.3 hours. In addition, 2500 pounds (1130 kilograms) of cargo would be loaded into the cargo bay and 10 passengers at the end of their space station tour-of-duty would board the orbiter for return to Earth.

Shortly after departing the space station, the flight crew would use the orbital maneuvering thrusters to perform a deorbit burn, then carefully orient the orbiter for reentry. It would enter the atmosphere moving at 25,912 feet (7900 meters) per second at an altitude of 400,000 feet (122,000 meters) and would slow to 20,000 feet (6100 meters) per second at an altitude of 200,000 feet (61,000 meters). At these speeds, the orbiter would compress the thin air in its path, causing severe aerodynamic heating.

GDC described the IPD Triamese Thermal Protection System (TPS) heat shield in greater detail than any other system. Mostly it would comprise overlapping metal "cover panels" backed by thermal insulation blankets. The company divided the TPS into windward (nose, belly, and leading edge) and leeward (everywhere else) sections.

The composition of the TPS cover panels and the composition and thickness of the insulation behind them would depend on many factors. These would include orbiter reentry angle, banking angle, potential for air cooling, location on the orbiter, and the existence of new development programs aimed at perfecting existing TPS materials or producing new ones. 

The majority of the panels would be mounted on posts attached to the propellant tanks, which were meant to serve as "primary structure." GDC modeled its tank design on that of the Saturn V S-II second stage, which it said was made up of "cylindrical integrated pressure tanks." These could carry structural loads while unpressurized except during launch and ascent. In areas where no propellant tanks were available — mainly over the cockpit and passenger compartment, the cargo bay, and the engine compartment — the panels would be mounted on posts attached to a "trapezoidal framework." 

For its IPD Triamese TPS calculations, the company assumed an entry angle no greater than 1°. This would yield skin temperatures ranging from 3950° Fahrenheit (F) (2180° Celsius — C) on the windward side of the orbiter nose to 700° F (370° C) on the leeward side of the fuselage 90 feet (27 meters) aft of the nose. 

Most of the IPD Triamese would be covered by TD Nickel-Chromium (TD Ni-Cr) panels capable of withstanding a reentry temperature of up to 2400° F (1315° C). TD Ni-Cr is a thorium oxide-coated alloy. The panels would measure just 0.01 inches (0.254 millimeters) thick. At that thickness, they would weigh 1.75 pounds (0.8 kilograms) per square foot (0.09 square meters). GDC estimated that the typical TD Ni-Cr panel could withstand 50 reentries before it would need to be replaced. 

The nose and rudder fin leading edges would create special TPS problems. GDC called a thorium oxide-coated tungsten nose cap a "representative" state-of-the-art system. This would, however, need to be replaced after every third flight, so the company called for accelerated development of new TPS materials. The rudder fin leading edges, which would be made of costly coated tantalum, would need to be replaced after every 10th flight. 

The insulation blankets behind the panels would comprise layers of Microquartz and Dynaflex, products of the Johns Manville Corporation. Microquartz, which would make up one-third of the thickness of the blanket when used with Dynaflex, would be made of silica microfibers. It could withstand temperatures up to 1600° F (870° C). Dynaflex, an aluminum oxide, silica, and chromium oxide microfiber material that could withstand temperatures up to 2800° F (1540° C), would make up the remaining two-thirds of the blanket thickness.

Insulation blanket thickness and composition would depend on location on the vehicle. It would, for example, consist of Microquartz and Dynaflex and measure 3.7 inches (9.4 centimeters) thick on the windward side of the cockpit/passenger compartment area. A layer of Microquartz alone just 0.8 inches (2 centimeters) thick would suffice on the leeward side beginning about 60 feet (18.3 meters) aft of the nose.

The orbiter would maneuver during hypersonic reentry using its rudder fin-mounted flaps and body flap-mounted elevons. Initial calculations showed that a 20° bank initiated at 400,000 feet (122,000 meters) would permit a landing up to 450 nautical miles (830 kilometers) off the orbital track while causing an average increase in surface temperature of only 40° F (23° C). More detailed calculations suggested a different approach: a 45° bank gradually reduced to 10° at 200,000 feet (61,000 meters), then gradually increased again to 45°.

GDC proposed that vehicle primary structure temperature be controlled through "detailed air injection" during flight. Vents in the fuselage would be opened during descent to admit air, then ducts would channel it to hot areas to keep the temperature below 200° F (93° C). The company calculated that failure to air-cool the IPD Triamese orbiter would allow heat to "soak" into the vehicle, driving primary structure temperature to a punishing 330° F (166° C) 50 minutes after landing.

Like the boosters during their return to KSC, the orbiter would slow to subsonic speed at an altitude of 15,000 feet (4570 meters). It would, however, reach that altitude nearer the KSC landing strip than would the boosters. The orbiter would then deploy its TF-34 jet engines and switchblade wings. Subsonic flight under jet power would last no more than 10 minutes. 

About 400 feet (120 meters) above the ground, the flight crew would lower the landing gear and perform a flare maneuver, raising the orbiter's nose so that its main landing gear would touch the runway first. The flight crew and passengers would feel a deceleration equal to two times Earth's gravity at touchdown. Landing would occur at a speed of 180 miles (290 kilometers) per hour; rollout would measure less than 10,000 feet (3050 meters) with switchblade wing flaps down and less than 13,000 feet (3960 meters) with flaps up.  Maximum landing weight was 135,300 pounds (61,370 kilograms).

Desk model of Triamese launch (left) and landing flare configurations. The landing flare configuration model displays switchblade wings (colored orange), one of two deployed TF-34 jet engines (colored silver), and tricycle landing gear. Image credit: National Air and Space Museum.

Immediately after landing, the orbiter would again enter the turnaround flow, joining the boosters with which it had launched a little more than a day before. GDC determined that, under normal circumstances, an IPD Triamese orbiter would require 810 person-hours of turnaround servicing, while a booster would need 490 person-hours. A normal orbiter turnaround could be completed in a week by two teams of 23 technicians working two eight-hour shifts. Flight data recorder analysis, mission planning, and payload preparation would need additional time. 

Occasional additional tasks would add to turnaround time. GDC envisioned a special engine inspection every six months and an annual three-day "calendar inspection," which would see technicians visually inspect the interior of the liquid oxygen and liquid hydrogen tanks along with all wiring and plumbing. Every two years, technicians would spend three weeks performing "progressive rework" maintenance, during which they would remove the entire TPS to allow a detailed inspection of all vehicle systems and system replacement and updating as necessary.

As the ILRV study continued into the Spring of 1969, NASA, often acting at the request of the USAF, imposed new requirements on its contractors. Most new requirements reflected an ongoing shift in reusable vehicle purpose away from low-cost space station resupply and crew rotation and toward general spaceflight cost savings. 

In April 1969, NASA asked the ILRV contractors to add a 15-foot-wide-by-60-foot-long (4.6-meter-wide-by-18.4-meter-long) payload bay to the orbiter component of their designs. The contractors were also directed to study designs that could place 50,000 pounds (22,680 kilograms) or 100,000 pounds (45,360 kilograms) of payload into low-Earth orbit. 

At about the same time, the space agency requested that they study orbiter missions independent of a station lasting up to 30 days. Such missions would, in effect, see the orbiter function as a short-term space station. This was an ill omen for NASA's ambitious space station aspirations. 

Adding a large payload bay and long-duration missions to the IPD Triamese orbiter undermined the cost-saving principle of boost element and orbiter element commonality. GDC sought to accommodate the new requirements within its Triamese proposal; for example, the company proposed clustering more than two booster elements around an expendable second stage attached to a large payload. By October 1969, however, it was clear that the Triamese concept's days were numbered. 

On 13 January 1970, NASA Administrator Paine announced that the Saturn V assembly line would be shut down permanently. AAP would, however, continue under the new name Skylab. The Apollo 20 Moon mission would be canceled so that its Saturn V could be stripped of its S-IVB third stage and put to work launching Skylab into Earth orbit. 

That same month, the ILRV study was redesignated Space Shuttle Phase A. On 28 January 1970, GDC teamed up with North American Rockwell (NAR) to compete jointly for a Space Shuttle Phase B contract, which they subsequently won. GDC applied its ILRV study experience to the design of a reusable Booster for an NAR reusable Orbiter.

Sources

"Togetherness," M. Getler, Aerospace Technology, 17 July 1967, p. 70.

"MOL Switch Forthcoming," Aerospace Technology, 1 January 1968, p. 3.

Memorandum, Douglas Lord, Deputy Director, Advanced Manned Missions Program, NASA Headquarters, to Maxime Faget, Manned Spacecraft Center, "Manned Spacecraft Center Revised FY 1967 Advanced Study Program," 10 April 1968.

"Pace of Post-Apollo Planning Rises," W. Normyle, Aviation Week & Space Technology, 3 February 1969, pp. 16.

"NASA Aims at 100-Man Station," W. Normyle, Aviation Week & Space Technology, 24 February 1969, pp. 16-17.

"Large Station May Emerge as 'Unwritten' U.S. Goal," W. Normyle, Aviation Week & Space Technology, 10 March 1969, pp. 103, 105, 109.

Triamese Reusable Launch Vehicle/Spacecraft Status Report II, Report No. GDC-DCB69-014, General Dynamics - Convair Division, 7 May 1969.

A Shuttle Chronology 1964-1973: Abstract Concepts to Letter Contracts, Volume I: Abstract Concepts to Engineering Data; Defining the Operational Potential of the Shuttle, Management Analysis Office, Administration Directorate, NASA Johnson Space Center, December 1988, pp. I-10 - I-15, I-81 - I-83, I-85, I-87 - I-95, I-101 - I-102, II-108 - II-110, II-138 - II-140, II-156, II-158 - II-159, II-166 - II-167, II-182 - II-184.

More Information

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

X-15: Lessons for Reusable Winged Spaceflight (1966)

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

Think Big: A 1970 Flight Schedule for NASA's 1969 Integrated Program Plan

McDonnell Douglas Phase B Space Station (1970)

Electromagnetic Launching as a Major Contribution to Space-Flight (1950)

Lunar electromagnetic launch track. Image credit: R. A. Smith/The British Interplanetary Society. 
In the 1968 novel 2001: A Space Odyssey, British author Arthur C. Clarke invoked an old spaceflight concept to begin a voyage to the Moon. Dr. Heywood Floyd, a space agency bureaucrat, boarded a winged, reusable Earth-to-orbit shuttle mounted horizontally atop a winged, reusable booster, which in turn was mounted horizontally on a "sled" riding on an electromagnetic track. 

The track used physical principles first studied in the late 18th century. It would not be too much of a stretch to think of it as a conventional rotary electric motor laid out flat, so that linear motion along a track replaced rotary motion about a shaft. The linear induction motor, as it is typically known today, was first studied as a means of launching aircraft by the Westinghouse Corporation in 1945.

Clarke's launch track activated as his booster's rocket motors ignited. The sled accelerated the booster/shuttle stack until the booster's wings began to provide lift. As it became airborne its rocket motors throttled up and the booster/shuttle stack began a rapid climb toward low-Earth orbit (LEO). 

Fifty-one years before the year 2001, Clarke outlined the limits of electromagnetic launching. In a paper in the November 1950 issue of the Journal of the British Interplanetary Society, he explained that a spacecraft could not attain Earth escape velocity of 11.2 kilometers (seven miles) per second on an electromagnetic launch track; as it gained speed, it would compress the air in front of it, subjecting it to aerodynamic heating sufficient to destroy it. He pointed out that the nose of the German V-2 missile had become "red-hot" at a speed of just 0.9 kilometers (0.6 miles) per second. 

Even if the thermal problem could be solved, launching a spacecraft bearing a crew on an electromagnetic track was unlikely to be practical. Humans cannot withstand high acceleration, so a piloted spacecraft would need a long launch track. Even if acceleration equal to 10 times the pull of gravity on Earth's surface were deemed acceptable, the electromagnetic launch track required to reach Earth escape velocity would need to be at least 600 kilometers (375 miles) long. 

Clarke solved these problems by moving the electromagnetic launch track to the Moon, where escape velocity is just 2.3 kilometers (1.4 miles) per second and there is no air. He also abandoned any thought of launching crews; his lunar electromagnetic launch track would be used to launch tanks full of rocket propellants manufactured from lunar surface material.

In 1950, no one knew the chemical composition of the lunar crust. Clarke assumed that the Moon could supply both hydrogen fuel and oxygen oxidizer. As it turned out, he was right, though this fact was not confirmed until the presence of water ice in permanently shadowed craters near the lunar poles was confirmed in the late 1990s. Water can be electrolyzed (split using electric current) to yield hydrogen and oxygen.

The power system for the electromagnetic launch track would depend on a flywheel that would be spun up gradually over hours or days using an electric motor driven by a nuclear or solar electric power source. Clarke calculated that a 50-metric-ton (55-ton) flywheel 4.4 meters (14.4 feet) in diameter spinning at 1200 rotations per minute could, if coupled to a properly designed overload-tolerant generator, provide an average power over two seconds of about one million kilowatts. This would be sufficient to accelerate a one-metric-ton (1.1-ton) cargo to a velocity of two kilometers (1.25 miles) per second. 

When electricity was applied to the track, acceleration would leap from zero to a "very high value" then decrease to 50 Earth gravities in just two seconds. This acceleration profile would dictate electromagnetic launch track length: it would stretch about three kilometers (1.9 miles) across the lunar surface. The track might be built on a level if no obstacles stood in its launch path; alternately, it could be built with a slight upward slope.

A propellant tank launched at two kilometers per second would slow to 0.78 kilometers per second as it reached the 3020-kilometer (1875-mile) high point (apoapsis) of its elliptical orbit about the Moon 2.5 hours after launch. If no further action were taken, it would reach the low point (periapsis) of its elliptical orbit five hours after launch traveling at two kilometers (1.25 miles) per second. Periapsis would occur on the lunar surface; in other words, the track-launched tank would crash. 

Clarke calculated that firing a small rocket motor at apoapsis to accelerate the propellant tank by 0.22 kilometers (0.14 miles) per second would result in a circular lunar orbit at apoapsis altitude. He assumed that the rocket motor would expend propellants equal to only 6% of the cargo weight, a quantity he deemed "trivial." A spacecraft could then rendezvous with the cargo and pump the propellants into its tanks.

A space station might corral propellant tanks after they arrived in circular lunar orbit, perhaps using small auxiliary vehicles. Spacecraft traveling in cislunar space could then rendezvous with the station to refuel. 

Alternately, the electromagnetic launch track might accelerate the tank to a slightly higher velocity, boosting it into an elongated elliptical orbit with a period of days. This would provide time for a spacecraft to rendezvous with its near apoapsis, where it would move very slowly. The empty propellant tank would be allowed to crash harmlessly on the lunar surface.

The electromagnetic launch track could, Clarke pointed out, place propellants mined and refined on the Moon into LEO at the cost of a 20% increase in launch velocity and increased guidance system complexity. The former could be achieved by lengthening the track and adding a "booster" flywheel/generator near its end.

If launching propellants to LEO were found to be feasible, then a cislunar spacecraft could reach LEO with empty tanks, refuel at a station, travel to lunar orbit, refuel at a station before landing, land, refuel at the lunar colony, climb to lunar orbit, refuel at a station, and depart lunar orbit for Earth orbit. Though operationally complex, this approach might simplify the design of cislunar spacecraft, since, as Clarke explained, none "need ever be designed for any mission more difficult than entry of a circular orbit round the Earth." 

Clarke cautioned that a lunar colony would need to be established and its "industrial potential" built up before lunar resources could be mined, refined, and fashioned into an electromagnetic launch track. "We are," he wrote, "rather in the position of trying to run a trans-Atlantic airline when there is no possibility of refueling. . .until we have drilled our own oil wells and set up our own refineries!"  

The time needed to establish a colony and industrial infrastructure on the Moon might, in fact, mean that technological breakthroughs would make launching propellants from the Moon using an electromagnetic track obsolete before it could begin. Clarke argued, nevertheless, that "it is. . .well to keep the Moon-based electromagnetic launcher in reserve as a solution of the long-term problems of spaceflight."

The image at the top of this post is Copyright © The British Interplanetary Society (https://bis-space.com) and is used by kind permission.

Sources

"Electromagnetic Launching as a Major Contribution to Space-Flight," Arthur C. Clarke, Journal of the British Interplanetary Society, Vol. 9, No. 6, November 1950, pp. 261-267.

The Exploration of the Moon, Arthur C. Clarke & R. A. Smith, Harper & Brothers, 1954, pp. 102-103.

2001: A Space Odyssey, Arthur C. Clarke, New American Library, 1999 (Millennial Edition), pp. 35-40.

More Information

Moon Suit: 1949

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

Humans on Mars in 1995! (1980-1981)

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

Could the Space Voyages in the Film and Novel "2001: A Space Odyssey" Really Happen? (Part 1)

Rendezvous Concept for Circumlunar Gemini (1965)

Graphic representation of a circumlunar journey. Image credit: Martin Marietta Corporation

On 18 August 1965, U.S. Representative Olin Teague of Texas, chair of the House Subcommittee on NASA Oversight and an ally of President Lyndon Baines Johnson, wrote a letter to NASA Administrator James Webb. "Much discussion is now taking place," the veteran Congressman wrote, "on the possibility of a circumlunar flight using a Gemini system prior to the Apollo lunar landing." Teague asked Webb for his opinion of the desirability of such a mission.

It was not the first time a piloted Gemini flight around the Moon on a free-return path — that is, without injection into lunar orbit — had been discussed. In late 1961, when Gemini was still called "Mercury Mark II" and NASA had yet to approve it as a formal program, a circumlunar flight had been proposed as one of its program objectives. The program was approved on 7 December 1961 and named Gemini the following month, but without the circumlunar flight. 

Gemini was envisioned as an experience-building bridge between relatively simple one-man Mercury flights and complex Apollo lunar landing missions. Use of rendezvous to accomplish President John F. Kennedy's objective of a man on the Moon by 1970 already seemed likely in early 1962. Rendezvous might take place in Earth orbit, lunar orbit, or both, and its challenges seemed daunting to many NASA planners. Gemini thus became seen as a rendezvous demonstrator.

In the spring of 1964, NASA Associate Administrator for Manned Space Flight George Mueller sought to pay McDonnell Aircraft Corporation, makers of the Mercury and Gemini spacecraft, to study a Gemini circumlunar mission. He saw the contractor study as an insurance policy; if Apollo suffered a major technical setback, or if the Russians looked set to carry out a piloted lunar flight, then circumlunar Gemini might salvage U.S. prestige. On 8 June 1964, however, NASA Associate Administrator Robert Seamans informed Mueller that Webb would authorize only in-house studies; NASA would not signal to its contractors a possible expansion or re-direction of Gemini.

Circumlunar Gemini came to Teague's attention 14 months later because astronaut Charles "Pete" Conrad, slated to serve as Gemini V pilot, was much taken with the concept. His enthusiasm led the Houston, Texas-based NASA Manned Spacecraft Center (MSC), home base of the astronauts, to study a circumlunar Gemini mission in collaboration with McDonnell and another major Gemini contractor — Martin Marietta Corporation, which manufactured the Gemini launch vehicle, the Gemini-Titan. Martin Marietta produced a report on the joint study in July 1965. 

The report lacked an MSC contract number and the Headquarters ban on NASA funding for contractor studies of circumlunar Gemini remained in effect. The companies apparently donated their time and expertise. 

The circumlunar Gemini mission described in the July 1965 report was scheduled to take place in June 1967, assuming a program go-ahead in September 1965. Use of existing or near-term planned hardware "building blocks" with minimal alteration would make the tight schedule possible. The building blocks included the Gemini-Titan and its larger cousin, the Titan IIIC launch vehicle, a modified Titan IIIC transtage upper stage, and a modified Gemini spacecraft. 

The Gemini-Titan was a Titan II Intercontinental Ballistic Missile modified to carry the two-person Gemini spacecraft. Modifications aimed mainly at improving safety. Among these were addition of backup systems and a Malfunction Detection System (MDS) that enabled the crew to monitor launch vehicle performance during ascent to low-Earth orbit. The Gemini-Titan, which launched from Pad 19 at Cape Canaveral Air Force Station (CCAFS), Florida, measured about 10 feet (three meters) in diameter and stood 107.65 feet (32.8 meters) tall with a Gemini spacecraft on top.

Gemini III launch, 23 March 1965. Image credit: NASA

By July 1965, the Gemini-Titan had flown four times. The Gemini I mission (8 April 1964) saw the two-stage rocket launch a simplified Gemini spacecraft into low-Earth orbit. Ballast replaced many missing Gemini spacecraft systems to give it a realistic weight and mass distribution. The spacecraft reentered and burned up as planned on 12 April 1964. 

Gemini II (19 January 1965) was a suborbital Gemini-Titan flight which ended with the first Gemini spacecraft splashdown and recovery. The third Gemini-Titan launched Gemini III, the first piloted Gemini spacecraft, on 23 March 1965. Mercury veteran Gus Grissom and rookie astronaut John Young orbited Earth three times before splashing down in the Atlantic Ocean.

Gemini IV (3-7 June 1965) saw James McDivitt and Ed White use their Gemini-Titan rocket for something other than ascent to low-Earth orbit. They tried unsuccessfully to approach and fly formation with its second stage, expending much more propellant than expected and, it appeared, confirming that the rendezvous maneuvers required in the Apollo Lunar-Orbit Rendezvous mission plan posed a significant challenge.

The first Titan IIIC rocket to fly stands on Launch Pad 40 at Cape Canaveral Air Force Station, 23 May 1965. Image credit: U.S. Air Force

The Titan IIIC launch vehicle was new in July 1965; its successful first launch had taken place on 18 June 1965. The 137-foot-tall (41.75-meter-tall) U.S. Air Force rocket comprised four stages. Stage 0 was a pair of Solid Rocket Motors (SRMs) that ignited simultaneously at liftoff. Each was about 10 feet (three meters) in diameter and 85 feet (25.9 meters) tall. 

The Titan IIIC SRMs were attached to the sides of a two-stage core closely resembling the Gemini-Titan rocket. The core stages, which burned Aerozine 50 fuel and nitrogen tetroxide oxidizer, were designated Stage 1 and Stage 2. Stage 1 ignited 105 seconds after liftoff, just before Stage 0 separation. It included a thermal shield to protect its engine assembly during Stage 0 operation, attachment points for Stage 0, and strengthened structure. It measured 10 feet (three meters) in diameter and 71 feet (21.6 meters) tall. Stage 2, 10 feet (three meters) in diameter and 37 feet (11.27 meters) tall, included strengthened structure and an extended interstage adapter to accommodate Stage 3.

A weather-beaten Titan IIIC transtage with a conical payload fairing arrives at NASA Johnson Space Center (JSC) in 2016. The old upper stage, destined for analysis by NASA orbital debris scientists, was transferred to NASA JSC after it was spotted in the aircraft "boneyard" at Davis-Monthan Air Force Base in Tucson, Arizona. Image credit: NASA
Stage 3, the fourth stage of the Titan IIIC, was the transtage, a restartable upper stage for boosting payloads from low-Earth orbit to higher orbits, including geosynchronous orbits. It measured about 10 feet (three meters) in diameter and 15 feet (4.6 meters) tall. Immediately after Stage 2 shutdown, retro-rockets ignited on Stage 2 to slow it, and Stage 3 slid along rails within the Stage 2 interstage adapter to ensure smooth separation.

The Titan IIIC transtage, with a pair of 8000-pound-thrust engines burning Aerozine 50 fuel and nitrogen tetroxide oxidizer, formed the basis of the most heavily modified circumlunar Gemini building block, the Modified Transtage (also called Transtage 2). The Martin Marietta/McDonnell/NASA MSC team sought to trim its weight so that it could boost an 8000-pound (3630-kilogram) Gemini spacecraft out of low-Earth orbit on a circumlunar path. They did this in part by relying on the Stage 3 Transtage attitude control system, telemetry system, and batteries. Removing these from the Modified Transtage reduced its weight.

They also added a Target Docking Adapter (TDA) borrowed from the Gemini Agena Target Vehicle (GATV), which at the time of their study had yet to fly. The GATV was, as its name implies, based on the Agena upper stage; in addition to giving Gemini crews a rendezvous and docking target, it would provide auxiliary propulsion for large orbit changes.

Gemini VI viewed from Gemini VII, 16 December 1965. Image credit: NASA
Cutaway of a Gemini spacecraft. Image credit: NASA

The final building block was, of course, the Gemini spacecraft. It comprised the Reentry Module and Adapter Module. The latter included the Equipment Module and the Retro Module. 

The Reentry Module included a pressurized cockpit with forward-facing windows, a blunt nose housing parachutes, attitude control thrusters, and rendezvous equipment, and a heat shield to protect it during Earth atmosphere reentry. The Gemini heat shield would be made sturdier and thicker to withstand the high-speed atmosphere reentry at the end of the circumlunar mission.

The Retro Module included solid-propellant deorbit rockets; these would be retained during the circumlunar Gemini mission to enable abort late in Gemini-Titan ascent to low-Earth orbit and to permit emergency reentry in the event that the mission could not depart low-Earth orbit. The Equipment Module, the broadest part of the Adapter Module, included the Orbit Attitude and Maneuvering System (OAMS) propulsion system and electricity-producing fuel cells.

The circumlunar Gemini flight program would begin with a Titan IIIC-launched heat shield qualification test without a crew in early February 1967. The Stage 3 transtage with attached stripped-down 5000-pound (2270-kilogram) Gemini would slide free of the Stage 2 stage at an altitude of about 100 nautical miles (185 kilometers) about 700 nautical miles (1295 kilometers) downrange from CCAFS. 

The transtage engine would fire for a short time, then the transtage-Gemini combination would coast to an altitude of about 150 nautical miles (280 kilometers) about 1500 nautical miles (2800 kilometers) downrange of the launch site. The transtage would then ignite for a second time, lofting the Gemini to an altitude of about 160 nautical miles (295 kilometers) about 2500 nautical miles (4630 kilometers) downrange before pitching over to drive the Gemini into the atmosphere. 

Transtage burnout and Gemini separation would take place about 3800 nautical miles (7040 kilometers) downrange at an altitude of about 120 nautical miles (220 kilometers). The modified Gemini would cast off its two-part Adapter Module and turn so its beefed-up heat shield faced in its direction of motion. Reentry at lunar-return speed of 36,000 feet (10,970 meters) per second would begin at 65 nautical miles (120 kilometers) of altitude about 4300 nautical miles (7960 kilometers) downrange, over the Atlantic Ocean near the space tracking facilities on Ascension Island. Splashdown and Reentry Module recovery would occur about 4600 nautical miles (8520 kilometers) downrange of CCAFS.

An Earth-orbital dress-rehearsal for the circumlunar flight would follow in mid-April 1967. The mission would test the rapid-fire launch, rendezvous, docking, and low-Earth orbit departure procedure McDonnell, Martin Marietta, and NASA MSC had selected for their circumlunar mission.

The test mission would begin with a Titan IIIC with a Modified Transtage and a Gemini-Titan with a Gemini spacecraft with two astronauts on board poised for launch on their respective pads. NASA would count down the two launches simultaneously. The Titan IIIC, with a shorter countdown, would reach launch (T) minus 30 seconds, then would be placed in a countdown hold. The Gemini-Titan would, meanwhile, count down to T minus six minutes and also be placed in a hold. 

After thorough system checkouts, the Gemini-Titan countdown would resume; 90 seconds later, the Titan IIIC countdown would restart, with T minus zero and liftoff taking place as the Gemini-Titan countdown reached T minus four minutes. Four minutes later, the Gemini-Titan countdown would reach T minus zero and liftoff would take place.

If all went as planned, the Gemini spacecraft would inject into an orbit 100 nautical miles (185 kilometers) above the Earth and separate from its Gemini-Titan second stage very near the Titan IIIC transtage and attached Modified Transtage. Ideally, rendezvous would occur at the moment of Gemini injection into low-Earth orbit. Launch dispersions were to be expected, however. The Martin Marietta/McDonnell/NASA MSC team was confident, however, that the Gemini spacecraft could inject into low-Earth orbit with its rendezvous target in range of its nose-mounted rendezvous radar.

The Gemini spacecraft and Titan IIIC transtage/Modified Transtage combination would orbit Earth every 90 minutes. One orbit after launch, the Gemini would be close enough to its target to begin a leisurely two-orbit "closure & docking" phase. Its slow pace would, it was hoped, conserve OAMS propellants. 

At the end of the closure & docking phase, the crew would insert their spacecraft's nose into the TDA on the front of the Modified Transtage. An electrical umbilical protruding from the nose would link to a receptacle in the TDA, enabling the astronauts to monitor and control the Modified Transtage. An external display panel on the TDA would also provide the astronauts with information on Modified Transtage systems.

A look inside the shrouds reveals a Titan IIIC transtage and, above it, the conceptual Modified Transtage for the circumlunar Gemini mission. A = streamlined payload fairing; B = Target Docking Adapter (TDA); C = TDA transition structure; D = payload fairing separation plane; E = Modified Transtage; F = Modified Transtage separation plane; G = Titan IIIC transtage/Stage 3; H = Titan IIIC transtage/Stage 3 separation plane. Image credit: Martin Marietta
Gemini docked with Modified Transtage. A = Gemini spacecraft; B = Target Docking Adapter (TDA) support structure; C = external status display panel visible to Gemini crew; D = TDA docking cone; E = Gemini electrical umbilical and TDA receptacle; F = TDA transition structure; G = Modified Transtage. Image credit: Martin Marietta.

Events would then occur rapidly. As the Gemini/Modified Transtage/Titan IIIC transtage stack orbited into the proper position to begin flight to the Moon, the crew would fire explosive bolts, severing links between the two transtages, then would ignite OAMS thrusters to pull the Modified Transtage clear of the Titan IIIC transtage. This would cause propellants in the Modified Transtage to settle toward its engines, permitting ignition.

With that, the April 1967 test would complete its main objectives. The astronauts on board the Gemini spacecraft would not ignite the Modified Transtage engines; instead, they would soon separate from the Modified Transtage and return to Earth. When time came for the actual circumlunar flight to begin in June 1967, however, the crew on board the docked Gemini would ignite the twin Modified Transtage engines within five minutes of separation from the Titan IIIC transtage, beginning the Trans-Lunar Injection (TLI) maneuver.

The Modified Transtage engines would fire for six minutes and 40 seconds, expending 22,565 pounds (10,235 kilograms) of propellants. At the start of the TLI burn, the crew would feel acceleration equal to 0.6 Earth gravities. Because they would face the Modified Transtage, they would feel as though they were falling out of their seats toward the Gemini spacecraft nose (straps would, of course, hold them firmly in place). Acceleration would mount up as the Modified Transtage expended its propellants and became lighter, reaching a maximum of five Earth gravities just before the engines shut down.

The astronauts would undock from the Modified Transtage, turn their Gemini spacecraft around, and fire the OAMS engines to move away. They would then settle in for a trip around the Moon.

The Martin Marietta/McDonnell/NASA MSC report contained few details on what the astronauts would do during their circumlunar voyage, apart from using the OAMS thrusters to carry out four course correction maneuvers. The first would take place during the period between three and 10 hours after TLI, the second and third  40,000 nautical miles (74,080 kilometers) before and after passing the Moon, respectively, and the fourth between five and 10 hours before Earth atmosphere reentry. Propulsive velocity change during the course-correction burns would total between 170 feet (51.8 meters) per second and 230 feet (70 meters) per second.

Other possible mission objectives included testing the worldwide tracking and communications system ahead of its use during Apollo lunar landing missions and lunar photography as the circumlunar Gemini passed over areas of the Moon lit by the Sun. The Martin Marietta/McDonnell/NASA MSC team estimated that about a third of the lunar farside hemisphere would be in sunlight as the spacecraft passed over it.

Flight time, maximum distance from Earth, and lunar passage distance depended on many factors and could be highly variable. For a circumlunar mission that would pass the Moon when it was near perigee and that would perform a splashdown near Cape Kennedy in daylight, the mission would last 143 hours (five days, 23 hours), would reach a distance of 221,700 miles (356,790 kilometers) from Earth, and would pass within between 660 nautical miles (1220 kilometers) and 1300 nautical miles (2410 kilometers) of the lunar surface. For a daylight splashdown near Hawaii when the Moon was near apogee, the equivalent numbers were 172 hours (seven days, four hours); 253,363 miles (407,748 kilometers); and between 800 nautical miles (1480 kilometers) and 1330 nautical miles (2460 kilometers).

Gordon Cooper (left) and Charles Conrad: the crew of Gemini V. Image credit: NASA
This post began with U.S. Representative Olin Teague's query to NASA Administrator James Webb. Three days after the date on Teague's letter, astronaut Pete Conrad, whose enthusiasm for a circumlunar Gemini flight had helped to inspire the Martin Marietta/McDonnell/NASA MSC study, reached orbit with Gordon Cooper on board Gemini V (21-29 August 1965). They doubled the voyage duration of Gemini IV and broke the world record for time in space (seven days, 23 hours). It was the first time the U.S. held that record — and it demonstrated that a human could live in space long enough to carry out a circumlunar voyage.

On 23 August 1965, while Gemini V orbited the Earth, Webb testified before the U.S. Senate Committee on Aeronautical and Space Sciences, chaired by Clinton P. Anderson of New Mexico, another ally of President Johnson. During his testimony, which marked the start of a three-day hearing on NASA's future, Webb reviewed work accomplished in the Apollo Program and sought support for an Apollo-derived post-Apollo space program. 

Without prompting, Webb briefly mentioned the circumlunar Gemini mission concept. If his aim was to elicit senatorial comment, he failed; the assembled Senators did not take the bait. The mission concept received no further mention during the three-day hearing.

On 10 September 1965, Webb responded to Teague. He explained that "insertion in our program of a circumlunar flight, using the Gemini system, would require major resources." Webb told Teague that "we are proceeding with many complex developmental, test, and operational efforts with too thin a margin of resources," adding that "if additional funds were available. . .it would be in the national interest to use these in the Apollo program." Webb included a copy of his Senate testimony with his letter.

At the end of September, Webb ordered his communications with Teague to be forwarded to Robert Gilruth, director of NASA MSC, Wernher von Braun, director of the NASA Marshall Space Flight Center, and Kurt Debus, director of NASA Kennedy Space Center, Florida. In an accompanying memorandum, Robert Freitag, director of Manned Space Flight Field Center Development at NASA Headquarters, explained that "this indicates NASA's position on possible circumlunar Gemini flights."

Sources

Rendezvous Concept for Circumlunar Flyby in 1967, Martin Marietta, July 1965

Letter, Olyn Teague to James Webb, 18 August 1965

National Goals for the Post-Apollo Period: Hearing on Alternative Goals for the National Space Program Following the Manned Lunar Landing, U.S. Senate Committee on Aeronautical and Space Sciences, 23-25 August 1965, U.S. Government Printing Office, 1965, p. 22

Letter with attachment, James Webb to Olyn Teague, 10 September 1965

Memorandum with attachment, Robert Freitag to various, 30 September 1965

Project Gemini: A Chronology, SP-4002, J. Grimwood, B. Hacker, and P. Vorzimmer, NASA, 1969, p. 153

On The Shoulders of Titans: A History of Project Gemini, SP-4203, B. Hacker and J. Grimwood, NASA, 1977, pp. 73-74, 200-201, 354

More Information

Around the Moon in 80 Hours (1958)

Gemini on the Moon (1961)

Space Station Gemini (1962)

The Spacewalks That Never Were: The Gemini Extravehicular Activity Planning Group (1965)