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.


"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 ( and is used by kind permission.


"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."


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)

Apollo to Mars & Venus: North American Aviation's 1965 Plan for Piloted Planetary Flybys in the 1970s

U.S. President Lyndon Baines Johnson (right) welcomes home astronauts Gus Grissom (center) and John Young (left) after their March 1965 Gemini III test flight. Earth-orbital Gemini was conceived as a means of bridging the yawning gap between "simple" one-man Mercury flights and complex three-man Apollo lunar-orbital and landing missions. Piloted Mars and Venus flybys based on Apollo technology might have played a Gemini-like role in the 1970s NASA program. Image credit: NASA.

A flyby is the simplest planetary exploration mission. We are accustomed today to seeing a flyby as a mission strictly reserved for automated spacecraft. In the early-to-mid 1960s, however, many within the NASA advance planning community believed that piloted flybys based on technology and techniques developed for the Apollo Moon program could enable U.S. astronauts to carry out effective exploration of Mars and Venus as early as the 1970s.

Much like robotic flybys, piloted flybys would limit themselves to small course-correction maneuvers after departing Earth. Robotic flyby spacecraft have no cause to return to Earth after passing their target or targets. Piloted flyby spacecraft, on the other hand, would swing past their target world or worlds on Sun-centered orbital paths that would intersect Earth, enabling their crews to return home.

The piloted flyby concept is usually attributed to Italian aeronautical engineer Gaetano Crocco, who in 1956 presented a paper describing a piloted Mars/Venus flyby. NASA-funded contractor work on piloted flybys began in 1962 with the Early Manned Planetary-Interplanetary Roundtrip Expeditions (EMPIRE) studies at NASA Marshall Space Flight Center (MSFC) in Huntsville, Alabama (see "More Information," below). EMPIRE also tasked its contractors with looking at piloted Mars and Venus orbiters. 

Crocco and NASA MSFC targeted their first missions for 1971, a date chosen with an eye toward limiting the time and propulsive energy required to reach Mars. In that year, Mars would be close to the Sun as Earth, nearer the Sun and moving faster, passed it. Because Mars has an eccentric orbit, this Earth-Mars geometry recurs only every 15 years or so. An opportunity for an Earth-Mars transfer as favorable would not occur again until 1988. 

The piloted flyby mission concept became increasingly attractive during 1964 and early 1965, when U.S. President Lyndon Baines Johnson made clear his vision of NASA's future after the Apollo Program. At that time, Apollo was expected to accomplish its first lunar landing during 1968. 

Johnson wanted Apollo lunar exploration to continue after the first successful landing, but mainly he wanted to see astronauts working on board Earth-orbiting laboratories derived from the Apollo spacecraft and Saturn rockets developed for the Moon program. These laboratories would, it was hoped, provide low-cost tangible benefits to American taxpayers through research in the fields of medicine, manufacturing processes, Earth resources discovery, agricultural monitoring, and advanced technology development. 

LBJ's vision of NASA's future made no mention of piloted Mars/Venus flybys based on Apollo's technological legacy. On the other hand, neither did it specifically forbid them.

A 15-month NASA MSFC in-house study begun in August 1963, shortly after EMPIRE ended, became, by its end, the first to examine the possibility of adapting Saturn rockets and Apollo spacecraft to piloted Mars and Venus flybys (see "More Information," below). Even before the NASA MSFC engineers completed their study in November 1964, NASA Headquarters and other NASA centers launched their own studies of Apollo-derived piloted flybys. The NASA Manned Spacecraft Center (MSC) in Houston, Texas, for example, completed an in-house study in February 1965. 

NASA MSC, which managed the Apollo Command and Service Module (CSM) spacecraft, contracted with North American Aviation (NAA), the CSM prime contractor, for a seven-month piloted flyby study that began on 1 October 1964. MSC then tacked on a two-month extension to NAA's contract so that the company could focus on enhancement of piloted flyby spacecraft long-term reliability through use of in-flight maintenance. NAA briefed MSC management on the results of its study in Houston on 18 June 1965.

The NAA study was significant in the evolution of piloted flyby planning because it was the first conducted by a major manufacturer of prospective piloted flyby hardware. In addition to the CSM, NAA was responsible for other Apollo hardware that might be adapted to a piloted flyby mission; specifically the Spacecraft-Launch Adapter (SLA) and the Saturn V rocket S-II second stage. 

This striking view of the Apollo 15 Command and Service Module (CSM) Endeavor in lunar orbit displays distinctive Apollo CSM features. These include the large Service Propulsion System (SPS) engine bell and the four-dish high-gain antenna (left), the slightly discolored housing for umbilicals linking the barrel-shaped Service Module (SM) with the silvery conical Command Module (CM), and the extended probe docking unit on the CSM's nose (right). Image credit: NASA.
This image displays the SLA shroud, the structural basis for NAA's piloted flyby Mission Module (MM), and, above that, the Apollo 11 CSM ColumbiaThe SLA, which linked the bottom of the CSM's silver-and-white SM to the top of the three-stage Saturn V S-IVB third stage, protected the Lunar Module Moon lander and the CSM's SPS engine bell during ascent through Earth's atmosphere. Please note launch gantry workers for scale. Image credit: NASA.
An NAA-built S-II Saturn V stage, with five J-2 engines, is slowly lowered into place atop an S-IC Saturn V first stage inside the immense Vehicle Assembly Building (VAB) at Kennedy Space Center, Florida. Please note workers for scale. Image credit: NASA

A brief aside is justified at this point: EMPIRE, the NASA MSFC and NASA MSC in-house studies, and the NAA study took place against the backdrop of Project Gemini. The two-seater Gemini spacecraft, the advanced cousin of NASA's first piloted spacecraft, one-man Mercury, was conceived as a training and biomedical research tool; it would provide astronauts, engineers, and flight controllers with experience in rendezvous, docking, and spacewalks, and would enable NASA doctors to certify that astronaut bodies could withstand roughly two-week Apollo lunar landing flights. 

In December 1961, NASA awarded McDonnell Aircraft Company, the Mercury prime contractor, the contract to build Gemini. Gemini I carried out a test flight without a crew a little more than two years later (8 April 1964). The first piloted Gemini, Gemini III, reached orbit with Gus Grissom and John Young on board on 23 March 1965. Project Gemini ended with its tenth piloted mission (Gemini XII) in November 1966.

Gemini served admirably as a bridge between Mercury and Apollo. Piloted flybys based on Apollo could, some felt, serve as a bridge between Apollo-derived Earth-orbiting space stations in the late 1960s/early 1970s and Mars/Venus orbiters and Mars landers in the late 1970s and 1980s. 

The early piloted flyby studies also took place against a backdrop of Mariner IV, which left Earth atop an Atlas-Agena rocket on 28 November 1964, nearly two months after the NAA study for NASA MSC began. When NAA briefed MSC managers, Mariner IV's planned 15 July 1965 Mars flyby was still a month away.

Few today would argue that robot probes need humans in close proximity to be able to accomplish their missions, but in the first years of the Space Age, it was different. Most robot probes failed. Those that succeeded delivered low-quality (though often tantalizing) data because they included relatively crude instruments and returned data at an agonizingly slow rate (Mariner IV was expected to return data at 8.3 bits per second; at that rate, 20 black-and-white images of the surface of Mars recorded on tape during its flyby would need a month to play back to Earth). 

Piloted flyby planners argued that a piloted flyby mission would be ideal for improving robot probe success rate and data quality. Astronauts could act as caretakers for a varied flock of probes they would release during the flyby. The probes would reach their target planet in tip-top condition. The piloted flyby spacecraft could act as a data relay, improving data rate. 

NAA added another argument: that robot probes released during a piloted flyby could be more sophisticated than those launched from Earth. Instruments and experiments could be made more complex (hence more prone to malfunction). Hitching a ride on a piloted spacecraft could also enhance probe flexibility; astronauts could, for example, direct an automated Mars lander to an intriguing site they had discovered through successive observations of increasing resolution made using a telescope on the piloted flyby spacecraft during approach to the planet.

NAA proposed a two-phase piloted flyby program. Phase I would see a piloted Venus flyby spacecraft launched in 1973 on a 415-day mission. During Phase II, which the company emphasized, a piloted Mars flyby spacecraft launched in 1975 would carry out a 700-day mission that would take it past Mars to the inner edge of the Asteroid Belt. In both phases, the piloted flyby spacecraft would comprise a modified CSM, a three-deck Mission Module (MM) containing living and working space for the crew, and a pressurized Probe Compartment (PC) bearing a cargo of automated probes tailored to their destination.

NAA envisioned that NASA would begin work on a formal piloted flyby Program Development Plan in mid-1966 and would award contracts to build the flyby spacecraft, robot probes, Saturn V rockets, and ground facilities with a "go-ahead" date of 1 July 1967. "Any slippage" in the go-ahead date, a "key milestone" in the piloted flyby program, would, NAA declared, "jeopardize the [19]73 and [19]75 launch window opportunities."

The company proposed a complex development, manufacture, and test program modeled on the one it was at the time following to build and test the CSM for Apollo lunar missions. Major Phase I milestones would include a test of a Command Module (CM) with a heat shield upgraded for high-speed reentry following a Venus flyby (June 1972), an Earth-orbital test of the complete Venus flyby spacecraft (August 1972), and a test of the Venus flyby spacecraft in solar orbit (December 1972-January 1973). 

Phase II milestones would include a test of the more robust Mars flyby CM heat shield and an Earth-orbital test of the Mars flyby spacecraft (December 1973-January 1974). Results from the Venus solar-orbital test could be extrapolated to the Mars flyby spacecraft, so no Mars solar-orbital test would be necessary.

NAA explained that the piloted Venus flyby would require at most two Saturn V launches, so could get by with the twin Complex 39 Saturn V launch pads (Pad 39A and Pad 39B) built for the Apollo lunar program at Kennedy Space Center, Florida. The piloted Mars flyby, on the other hand, might require as many as four Saturn V launches in rapid succession, so NASA would need to build two new Saturn V pads. Pad 39C and Pad 39D would be ready in August 1974.

The Phase I Venus flyby mission would leave Earth orbit on 30 October 1973. The Phase II Mars flyby would depart on 5 September 1975. Only the Mars flyby will be described in detail here, in keeping with NAA's emphasis on Phase II. 

Piloted flyby spacecraft in Earth-orbital configuration. A = piloted flyby Command Module (CM); B = piloted flyby Service Module (SM); Mission Module (MM); D = Probe Compartment (PC); E = docking adapter linking PC to S-IIB orbital launch stage; F = S-IIB orbital launch stage; G = adapter for linking S-IIB stage to two-stage Saturn V launch vehicle (discarded before launch from Earth orbit). Image credit: North American Aviation/NASA.

NAA assumed that its flyby missions would be boosted from Earth orbit by an S-IIB orbital launch stage (see "More Information," below), a modified version of NAA-built S-II, the second stage of the Apollo Saturn V. The piloted flyby spacecraft and the S-IIB would launch separately on two-stage Saturn V rockets and dock in a 262-nautical-mile-high (485-kilometer-high) assembly orbit. 

The Mars flyby S-IIB stage would, if loaded with the necessary propellants, be too heavy for the two-stage Saturn V to deliver to assembly orbit. NAA proposed that the S-IIB be launched with a full load of liquid hydrogen fuel and an empty liquid oxygen tank. One or two Saturn V-launched liquid oxygen tankers would then dock with the S-IIB to fill the oxygen tank in orbit. After the oxygen tank was full, the tanker would withdraw and deorbit itself over a remote ocean area. The Mars flyby spacecraft and S-IIB stage would dock, then the latter would ignite its J-2 engines to begin the journey to Mars.

Cutaway of four-person flyby Command Module (CM). A = heat shield; B1 = forward middle crew couch; B2 = right crew couch; B3 = after middle crew couch; C = Apollo-type probe docking unit; D = housing for life support and data umbilicals linking CM to Service Module (SM); E = mercury-rankine isotopic power system; F = housing for isotopic power system cooling and electricity umbilicals linking CM to SM. Image credit; North American Aviation/NASA.

Citing the many responsibilities of the crew during Mars close passage, NAA argued for a four-person Mars flyby crew. To make room for a fourth astronaut in the Mars flyby mission CM, the center launch-and-reentry couch would be relocated forward of its Apollo CM position, placing it closer to the main display and control console. A new fourth couch would be mounted on the aft interior bulkhead about two feet behind the relocated center couch. The left-hand couch and the right-hand couch would remain in their Apollo CM positions. 

NAA reminded its MSC audience that the Mars flyby CSM would be called upon to support its crew for a much shorter period of time than would the Apollo CSM. The flyby crew would reach and depart Earth-orbit on board the flyby CSM, return to Earth in the flyby CM in the event of an abort during the hour immediately after Earth-orbit departure, briefly power up the flyby CSM and fire its center engine during the mission's anticipated eight modest course corrections, and return to Earth's surface in the flyby CM at the end of their mission. The company estimated that the flyby astronauts would inhabit the flyby CM cabin for no more than 72 hours at a stretch, not the 10 or more days of a lunar mission.

Image credit: NASA/DSFPortree.

The most obvious external modification to the CSM for NAA's piloted Mars/Venus flyby missions was replacement of the Apollo CSM's single Service Propulsion System (SPS) main engine with three modified Lunar Module (LM) descent engines, each with independent propellant tanks and plumbing. Two half-cone housings added to the sides of the Service Module (SM) would provide room for the two outboard engines. 

Any single flyby CSM engine could perform all necessary flyby mission maneuvers, NAA declared. If all three rocket engines remained functional throughout the flyby mission, however, the middle engine would be used to perform course corrections and the two outboard engines would together carry out the retro burn at the end of the Mars or Venus flyby mission. 

An abort at the start of the Earth-Mars transfer, during the hour following Earth-orbit departure, would burn propellants which would, in a successful mission, be used to perform course corrections and to slow the Mars flyby CSM ahead of Earth-atmosphere reentry. The abort burn would expend nearly all of the Mars flyby CSM's propellants.

Assuming that no abort were necessary, the flyby astronauts would cast off a cylindrical two-part adapter linking their CSM to the top of the MM. They would then move the CSM away using reaction control thrusters, turn the CSM end for end to face the MM, and dock with an Apollo-type drogue docking unit on top of the MM. The crew would then shut down the CSM and transfer to the 5600-cubic-foot MM, their home for the next 700 days.

The MM drogue unit would be inset within a housing that would, after docking, encase the conical CM. If NASA opted for a weightless environment for its piloted flyby crews, the housing would shield the CM from meteoroid damage. If, on the other hand, NASA opted for an artificial-gravity environment, the housing would be split into two parts. The upper part would latch onto the sides of the CM below its windows; the lower part, firmly attached to the MM, would contain cable reels. 

Transition from zero-gravity configuration to artificial-gravity configuration. Image credit: North American Aviation/NASA.

The crew in the MM would spin up the piloted flyby spacecraft using thrusters in the PC. After a gentle nudge from the thrusters, they would unlatch connectors linking the upper and lower parts of the housing and begin to pay out the cables. The CSM, linked to the cables by the "collar" formed by the upper housing, would move away from the MM/PC combination. This would slow the rate of spin about the flyby spacecraft center of gravity, which would in turn reduce tension in the cables, raising the possibility of tangling. 

The crew would, however, continue to fire the thrusters in brief bursts, slowly increasing the spin rate and keeping cable tension constant. When the cables reached full extension, the CSM and MM/PC would be 158 feet (48.1 meters) apart, completing four rotations per minute. This would provide the crew in the MM with acceleration that they would feel as gravity roughly equal to the pull of gravity on Mars (0.4 G). Providing the crew with Mars-level gravity complemented the flyby mission biomedical research program; data on human response to Mars-level gravity would clear the way for long stays on the surface of Mars in the 1980s.

The piloted Mars flyby spacecraft would spin for 660 days of its 700-day voyage. The 40-day non-spinning period would include course-correction rocket burns using the center CSM engine at 73 days, 139 days, 260 days, 472 days, and 550 days, plus an unspecified period surrounding the Mars flyby on 2 February 1976, 150 days into the mission, during which spin would be stopped to facilitate Mars observations and release of robot probes. Spin-down would require a reversal of the spin-up process; the crew would activate the cable reels to slowly retract the CSM while burst-firing thrusters in the PC to decrease the spin rate gradually.

After spin-down and spin-up, the flyby crew would need to reorient their main communications link with Earth, the 13.1-foot-diameter (4-meter-diameter) high-gain dish antenna mounted on a boom on the PC. The high-gain was designed to spin at four rotations per minute in the direction opposite the piloted flyby spacecraft's spin, enabling it to maintain a constant lock on Earth. During periods when the flyby spacecraft did not spin, the high-gain rotation motors would make slight adjustments to its orientation to maintain a steady lock on Earth.

A simplified view of the major components of the flyby CSM's electrical power system. A = the mercury-rankine isotopic power system; B = umbilicals for circulating cooling fluid from the power system in the CM to the radiator panels on the SM and back again; C = redundant radiator panels for cooling the mercury-rankine isotopic power system. Image credit: North American Aviation/NASA.

NAA determined that using the CSM as an artificial-gravity counterweight created an opportunity. The company proposed that the CM include a compact 1370-pound (620-kilogram) plutonium-fueled mercury-rankine isotopic power system capable of generating four kilowatts of continuous electricity for the flyby CSM, the MM, and the PC. If it was to be ready in time for the 1975 Mars flyby mission, NAA estimated, development of the isotopic system would need to start in July 1965 — that is, less than two weeks after the company briefed NASA MSC.

Putting the isotopic system in the Mars flyby CM would place it at a distance from the crew throughout most of the mission, so would expose them to a negligible radiation dose. Special-purpose shielding and water for evaporative cooling of the isotopic system after CM separation from the SM just before Earth atmosphere reentry would shield the flyby astronauts from radiation while they were inside the CM. NAA was confident that the Mars flyby crew would receive an acceptably low cumulative radiation dose from the isotopic system during the brief time they rode in the Mars flyby CM. 

Umbilical hoses would link the isotopic system in the flyby CM to redundant radiator panels on the SM's hull. Liquid metal (potassium-sodium) coolant would flow through the isotopic system, hoses, and radiator panels in a continuous loop. NAA envisioned using the same cooling loops for CSM life support system cooling.

NAA's chief justification for reliance on an isotopic source had to do with the Mars flyby mission's maximum distance from the Sun. The spacecraft would race past Mars on a low-energy path that would take it to the inner edge of the Asteroid Belt, more than 200 million miles (320 million kilometers) from the Sun. It would then fall back toward the Sun and intersect Earth. The solar arrays required to generate four kilowatts of electricity continuously at that distance would be prohibitively large and heavy. Their extent would make them prime targets for marauding meteoroids, which were expected to become a significant hazard as the spacecraft skirted the Asteroid Belt.

The Venus flyby spacecraft, by contrast, could rely on an ample solar energy supply and, it was expected, would contend with a meteoroid population less dense than found at Earth. NAA assumed that a 525-pound (240-kilogram) solar-cell power system would be adequate to power the Venus flyby spacecraft.

The Mission Module (MM) with major components indicated by letters. A = crew transfer tunnel leading to the Probe Compartment (PC); B = hatch and retractable ladder; C = Probe Compartment; D = shelter/control center (lower deck); E = centrifuge; F = middle deck (main living and working area); G = sleep area; H = crew transfer tunnel linking the drogue docking unit to the top and middle decks; I = gaseous oxygen tank and life support equipment; J = liquid nitrogen tank; K = liquid oxygen tank; L = docking collar; M = drogue docking unit; N = two-part adapter linking flyby CSM and MM. Image credit: North American Aviation/NASA.

The four-segment SLA, the NAA-built adapter that linked the Apollo CSM to the top of the Apollo Saturn V S-IVB third stage, would form the structural basis for the piloted flyby MM, the crew's home and workplace during interplanetary travel. NAA did not design an MM specifically for its piloted flyby study. Instead, it tapped the Apollo Orbital Research Laboratory, a 1962-1963 NAA concept for a small space station based on the SLA structure. 

The tapered MM would include three decks with a total of 800 cubic feet (22.65 cubic meters) of open space per crewmember. The top deck, smallest of the three, would include at its center a crew transfer tunnel, which would lead from the drogue docking unit atop the MM to the ceiling of the middle deck. Liquid oxygen and liquid nitrogen tanks would surround the upper part of the transfer tunnel just above the top deck ceiling. The top deck, accessible through an opening in the side of the tunnel, would contain sleeping, medical, and hygiene facilities, as well as life support equipment and a large tank of high-pressure gaseous oxygen.

NAA described its MM life support system in some detail. During the first year of the 700-day Mars flyby mission, the crew would breath oxygen and nitrogen stored in dense, super-cold liquid form; they would then switch to oxygen stored as gas. The large tank on the top level of the MM could completely pressurize the module six times over the course of the mission; this might become necessary to flush out built-up trace gases outgassed from furnishings and produced during soldering, food preparation, and other processes.

The crew would take in oxygen and exhale carbon dioxide. NAA proposed to split the carbon dioxide to recover oxygen using the Bosch reaction, which uses hydrogen and produces carbon and water. The water would then be electrolyzed to yield hydrogen and oxygen. NAA calculated that, assuming 10 pounds of air leakage per day, the piloted Mars flyby mission would need to carry a total of 11,035 pounds (5005 kilograms) of oxygen. 

Ladder rungs in the transfer tunnel would continue as a ladder on the middle deck, the crew's main living area. The middle deck would include the galley, multipurpose table, equipment for making repairs and performing data reduction, and portholes with provisions for securely mounting cameras and other instruments. 

The bottom deck, widest of the three, was probably the most interesting. It would contain a centrifuge for subjecting astronauts to acceleration equal to the pull of gravity on Earth. The centrifuge would include two seats and two storage cabinets which between them would hold more than 900 pounds of Mars flyby spacecraft spare parts. The cabinets would serve as counterweights, stabilizing the centrifuge.

The centrifuge would spin around the "storm cellar" shelter/control center, a 600-cubic-foot (17-cubic-meter) bell-shaped compartment that could be sealed off from the rest of the flyby spacecraft. It could support the four-person crew for up to four days without resupply, allowing them to safely ride out solar flares. To save weight, the shelter/control center would contain little special-purpose radiation shielding, relying instead on its central location on the flyby spacecraft's widest deck and the bulk of equipment — centrifuge, spare parts cabinets, and control consoles for operating MM/PC systems — surrounding it.

NAA described a regular day in the Mars flyby crew's voyage. The crew would sleep for six hours, work for eight hours, grab a 1.5-hour nap, then work again for 8.5 hours. Work periods would be interspersed with four 20-minute periods set aside for eating and 50 minutes total for personal hygiene. The company expected that on average each crewmember would spend about 6.5 hours per day on MM and probe maintenance, and 2.5 hours per day advancing the flyby mission science program.

Exercise would count toward work time: in the hope of counteracting the effects of life in Mars gravity, NAA scheduled 1 hour of light exercise, 30 minutes of medium exercise for biomedical monitoring, and 30 minutes of heavy exercise. Crewmembers would spend one hour per day riding the centrifuge. 

Probe Compartment (PC) exterior. A1 = side view of high-gain antenna; A2 = partial front/rear view of high-gain antenna; A3 = high-gain antenna dish folded prior to deployment; B = magnetometer boom (side and aft views); C = 40-inch (one-meter) telescope (side and aft views); D = cutaway of PC showing interior structure; E = PC aft pressure hull; F1 = deployment panel for Soft-Lander Probe (SLP) 2; F2 = deployment panel for SLP 1; F3 = deployment panel for Orbiting Environment Monitor (OEM) and Orbiting Astronomical (OAT) probe; F4 = deployment panel for Hard-Landing Probes (HLPs); F5 = deployment panel for Parachuted Atmosphere Probes (PAPs) 1, 2, and 3; F6 = deployment panel for PAPs 4, 5, and 6. Image credit: North American Aviation/NASA.
Cutaway of Probe Compartment showing probes. A = spin-up/de-spin motors; B = spin-up/de-spin propellant tank; C = probe propellant tanks; D = Soft-Lander Probe (SLP) 2; E = Parachuted Atmosphere Probes (PAPs) 4, 5, and 6; F =  Hard-Landing Probes; G = SLP 2; H = Orbiting Astronomical (OAT) probe; I = Orbiting Environment Monitor (OEM) (PAPs 1, 2, and 3 behind). Image credit: North American Aviation/NASA.

A hatch in the middle of the shelter/control center floor would lead to a crew transfer tunnel. The tunnel would in turn lead to the PC, which would contain 15 probes of six types with a combined weight of between 6000 and 12,000 pounds (2720 and 5440 kilograms), telescoping launchers, and tanks of "sterilization gas" of unspecified composition for killing Earth microbes ahead of probe launch.
The PC would include a pair of probe maintenance stations which would between them feature a folding work bench, displays for monitoring probe health, and 94 cubic feet of storage including 65 cubic feet of probe spare parts storage. In addition, it would carry spherical tanks containing unspecified propellants for the two Mars orbiters.

The orbiters, designated the Orbiting Environment Monitor (OEM) and the Orbiting Astronomical (OAT) probe, would be the first of the carefully tended probes to be launched. Each would include a two-stage propulsion system. The first stage was intended to deliver the probe to Mars ahead of the piloted flyby spacecraft; the second would slow it so that the planet's gravity could capture it into Mars orbit. 

Minus their two rocket stages, they would take the form of 60-inch-by-125-inch (152-centimeter-by-318-centimeter) drums. The OEM would weigh 3900 pounds (1770 kilograms) and the OAT, 4390 pounds (1990 kilograms). Each would include an extendable solar array/instrument platform mounted on pivoting arms. They were expected to operate independent of the piloted flyby spacecraft for up to 180 days. In addition to gathering data using their own instruments, they would relay data from two Soft-Lander Probes (SLPs) on Mars.

Probes meant to enter the martian atmosphere would all have "blunt" shapes; NAA hoped that this would cause them to decelerate rapidly in the upper martian atmosphere, allowing them to descend slowly toward the surface, gathering data for as long as they could. Most would be shaped like the Apollo CM. Five Parachuted Atmosphere Probes (PAPs) were the exception; each would take the form of a 24-inch (61-centimeter), 160-pound (72.6-kilogram) sphere. 

The PAPs were intended to operate for just 200 seconds before they crashed into the surface of Mars. Only the six Hard-Landing Probes (HLPs) had shorter planned useful lives; each 47-inch (120-centimeter), 150-pound (68-kilogram) HLP would return data for just 100 seconds.

SLP 1 was the largest lander; it was it would measure 12.8 feet (3.9 meters) in diameter and would weigh 1870 pounds (848 kilograms). Meant to operate for 180 days, it would carry a variety of scientific instruments, including an Automated Biological Laboratory (ABL). The ABL would, as its name implies, gather samples of its surroundings to seek out biology. 

In 1964-1965, many scientists expected to find microbial life on Mars; not a few anticipated that higher forms, perhaps resembling moss, lichen, or even lithops ("living stones") or cacti, might occur. A few scientists — possibly not the greatest logicians in the scientific community — expected that plants naturally meant that animals should exist to eat them. The ABL, which was proposed in many forms in the early 1960s, would carry a complex payload of life-detection instruments intended to anticipate all of these possibilities.

SLP 2 would be less that half as heavy as SLP 1 (just 840 pounds/381 kilograms), yet would encompass within its 9.3-foot (2.85-meter) diameter a variety of intriguing (and poorly described) subprobes. These would include three "projectile" probes, three balloon probes, and a "TV probe."  

In addition to the probes, the PC would carry mounted on its exterior a 40-inch (one-meter) telescope and a rear-pointing magnetometer boom. The telescope, which would be used for many planetary science and astronomy objectives during the 700-day mission, could be steered and slewed to track on Mars during the flyby. This would avoid photographic image smearing. NAA envisioned equipping the telescope with folding, steerable mirrors to expand its field of view during the flyby, enabling it to track on the surface below the speeding spacecraft and on the horizon simultaneously.

NAA listed 28 Mars flyby mission primary science and engineering objectives, most of which aimed to prepare the way for more advanced piloted Mars missions in the 1980s. Scientific exploration was, of course, not to be neglected during the flyby mission, but the company took pains to stress that science would not become the chief mission emphasis until NASA conducted orbiter and landing missions. 

On 2 February 1976, 150 days after Earth departure, NAA's piloted Mars flyby spacecraft would reach its target. The company's representatives told MSC managers that the 32 hours surrounding "periplanet" — as it called closest approach to the surface of Mars — would be the mission's "pay-off phase." 

The top-priority objective would be to collect photographic data required to make detailed Mars maps. The crew would observe and photograph Mars using the telescope and a 35-mm film camera with a "turret" of different lenses mounted on a flyby spacecraft porthole. 

Mars maps in 1965 included few surface features beyond the largest light and dark areas. They were largely based on photographic plates taken from Earth using large telescopes and sketches made by astronomers peering through telescope eyepieces. Most still contained at least a hint of the "canals" first noted by Giovanni Schiaparelli in 1877 and popularized by Percival Lowell beginning in the 1890s. 

The crew would monitor and take data from the robot probes, which they would release at carefully determined times to ensure that they would reach targets selected on the basis of telescope observations made during approach to the planet. Radio signals from the probes would be received through an antenna attached to the flyby CSM in place of the Apollo CSM's four-dish high-gain antenna. 

The crew would, as might be expected, alter their regular daily schedule during the flyby. Sleep would be reduced by 1.5 hours per crewmember per day, eating time would be cut in half, and exercise and biomedical monitoring would be eliminated. NAA allotted 4.5 hours per crewmember for probe monitoring, two hours for non-probe science using the telescope and 35-mm camera, and three hours for "unscheduled" observations (The company suggested, for example, that the astronauts might wish to sketch what they saw on Mars).

NAA plotted the ground track the flyby spacecraft would follow from 24.8 hours (one martian day) before periplanet to 24.8 hours after periplanet. At the start of that period, an entire martian hemisphere would be in view centered on the nondescript light-colored region Aethiopis, about 10° north of the equator. The ground track would then run westward, passing over dark-colored Syrtis Major and light-colored Aeria. 

At 18.6 hours before periplanet, the flyby spacecraft would enter the martian "sphere of influence" and would begin to accelerate under the pull of martian gravity. Between that time and 12 hours before periplanet, it would pass over the light-colored regions Eden, Chryse, and Xanthe, north of dark-hued Sabaeus Sinus, Meridiani Sinus, and Margaritifer Sinus. 

Twelve hours before periplanet, the ground track would pass through little-hued Candor. By that time, the flyby spacecraft would be close enough to Mars that the field of view outside the portholes would take in a region between about 55° north and 35° south latitude and from 30° west to 140° west longitude. 

With the flyby spacecraft moving ever faster, the ground track would sweep west over light-hued Tharsis and Amazonis south of the enigmatic bright spot Nix Olympica. Six hours ahead of periplanet, the field of view would take in Amazonis between 30° north and 10° south latitude. Features a kilometer across would become readily visible through the telescope. 

In the last three hours before periplanet, the ground track would sweep south of mysterious Elysium. On Mars maps available in 1965, Elysium was a light-hued pentagon bounded by five diffuse canals. 

Finally, the track would turn northwest and sweep across light-colored Arabia. Elevated features on Mars would by then show west-pointing shadows; the piloted flyby spacecraft would race toward night, and behind it the Sun would sink rapidly toward the planet's limb. 

Minutes before periplanet, with the ground track passing just south of Cydonia, the Sun would set. Periplanet would take place in faint twilight, with the surface cloaked in blackness, at an altitude of 189 miles (305 kilometers). The piloted flyby spacecraft would then begin to move away from Mars. 

The crew would use the flyby CSM center engine to perform a small course correction immediately after periplanet. The maneuver would compensate for the effects on the spacecraft's course of any irregularities in the martian gravitational field. Performing the correction close to Mars would reduce the quantity of propellants required to carry it out.

Some of the Mars feature names mentioned above will sound familiar, for many were preserved, sometimes in slightly altered form, after U.S. robotic spacecraft mapped Mars. Candor, for example, lent its name to a section of Valles Marineris, the great equatorial rift and canyon system imaged by the Mariner 9 orbiter in 1971-1972. Meridiani Sinus was renamed Terra Meridiani; it became the landing area for the Opportunity rover in 2004. Chryse is now Chryse Planitia; the Viking 1 lander performed the first successful Mars soft-landing there on 20 July 1976. The name Tharsis was applied to a vast volcanic bulge atop which rise four shield volcanoes; one of these, Olympus Mons, is the largest volcano known in the Solar System. 

The Syrtis Major hemisphere of Mars. Syrtis Major Planum is the dark feature at the center of the image; the light area to its left is Arabia Terra and the dark area on the limb at left is Meridiani Terra. Image credit: NASA.
The Tharsis hemisphere of Mars. Patches of cloud mark the four great shield volcanoes; Olympus Mons is above and to the left of center. Western Valles Marineris is on the limb at center right. Image credit: NASA.

All of these surface features would be visible to the four astronauts during the Mars flyby. NAA assumed that only robotic precursors of minimal capability would precede them to Mars, so they would become the first humans to glimpse its wonders.

NAA compared its piloted flyby with planned robotic Mars missions. The company told NASA MSC managers that the Voyager probe proposed for launch in 1975 (not to be confused with Voyager outer planets probes launched in 1978-1979) would transmit data at a rate of between 100 and 350 bits per second. The piloted flyby mission, in stark contrast, would return 2000 bits per second and would deliver rolls and cassettes of high-resolution film to cartographers and researchers on Earth. NAA declared that its analysis had shown that "types, ranges, accuracies, and quantities of data obtained [by a piloted Mars flyby mission] should exceed (by orders of magnitude in some cases) that which could be returned to Earth with equivalent instruments on unmanned systems."

With Mars flyby activities tapering off, the crew would return to their regular daily schedule and commence the 550-day voyage home. They would begin Mars data analysis and, as they skirted the Asteroid Belt, observe any nearby asteroids using their telescope. 

The crew would also pay close attention to the health of their spacecraft's systems during the long trip home. They would have at hand tools and carefully selected spares to perform repairs. These would be available in part as a result of the two-month study of piloted Mars flyby spacecraft systems reliability NASA MSC added to NAA's original study task. 

The company estimated that 57% of piloted Mars flyby spacecraft subsystems — which included life support, power, propulsion, guidance, communications, and data handling — could be provided by 164 hardware "assemblies" designed for Apollo lunar missions. Another 22% (63 assemblies) could take the form of modified Apollo hardware, and 15% (44 assembles) could comprise hardware borrowed from other programs, such as the U.S. Air Force Manned Orbiting Laboratory. 

This meant that 94% of piloted Mars flyby hardware would have a test record and failure history by the time the piloted Mars flyby mission left Earth in 1975 even if the Phase I Venus flyby did not fly in 1973. The remaining 6% (just 17 assemblies) would require new development and testing.

Based on existing Apollo reliability predictions, NAA calculated that from six to 85 failures would occur during the 1975 piloted Mars flyby mission. Most would occur in subsystems that could be repaired or replaced by the crew. Those assemblies that could not be repaired or replaced — for example, the large thermal radiator on the outside of the MM — could be modified during the design phase to avoid failure or backed up by a redundant system.

NAA became concerned that some subsystems would take so long to repair that the crew could be harmed by the malfunction before they could finish. Analysis showed, however, that no repair time would exceed allowed downtime. A failed cabin heat control system, for example, could be repaired in an hour but would need from eight to 24 hrs to create a problem sufficiently serious that it would harm the crew. 

The company found that up to 185 spares weighing about 900 pounds would be required as insurance against all possible failures. Of course, very few were likely to be used. Repair time spread over the mission would amount to only about 15 minutes per day. 

Return to Earth would occur on 5 August 1977. As Earth grew large outside the portholes, the flyby spacecraft crew would prepare to abandon their home of 700 days. They would reel in the CSM for the last time and load it with film and other data. About two hours before planned reentry they would separate the CSM from the drogue docking unit and the artificial-gravity collar on the MM and back away. 

The crew would orient the Mars flyby CSM so its three engines pointed in its direction of travel and, 30 minutes before planned reentry, would ignite the two outboard engines. Flyby mission Earth-return speed would depend on many factors: for example, a close Mars flyby typically would mean a fast Earth-atmosphere reentry.

The Apollo CM was designed to reentry Earth's atmosphere at 36,000 feet (10,970 meters) per second. NAA told MSC that the CM's bowl-shaped heat shield could, in theory, be beefed up to withstand reentry at 52,000 feet (15,850 meters) per second. The company argued, however, that "engineering conservatism" made such high-speed reentries unattractive. Hence the retro burn, which would slash reentry velocity to no more than 45,000 (13,715 meters) feet per second. NAA told NASA MSC that the Apollo CM heat shield would need only modest modifications to withstand reentry at that velocity.

NAA reported that, at launch from Earth, the Apollo CSM would have a mass of 57,690 pounds (26,170 kilograms). Hypergolic (ignite-on-contact) Hydrazine/nitrogen tetroxide propellants would account for 37,360 pounds (16,950 kilograms) of that total. The hefty Mars flyby CSM would have a mass of 73,080 pounds (33,150 kilograms) of which 44,770 pounds (20,310 kilograms) would constitute propellants for course corrections and the reentry retro burn. 

During the retro burn, the outboard engines would fire for up to 29 minutes to slow the flyby CSM. The flyby SM would then separate, exposing the CM's modestly uprated heat shield and depriving the isotopic power system of its heat radiators (it would switch to its temporary water boil-off cooling system). During passage through Earth's atmosphere, the heat shield might attain a temperature of 5000° F (2760° C). 

NAA anticipated that the Mars flyby CM would parachute to a land landing. Modifications to the shock absorbers in the crew couches would protect the astronauts from injury as the CM bumped to a stop on Earth. Soon after landing, the isotopic power system would boil off the last of its cooling water; hence, linking it to an ground-supplied auxiliary cooling system would be assigned nearly as high a priority as removing the astronauts from the CM.

Direct Venus flyby and "in transit" assembly Mars flyby Saturn V launch configurations. A = Launch Escape System tower; B = piloted flyby Command and Service Module (CSM); C = Mission Module (MM); D = Probe Compartment (PC); E = Saturn V S-IVB stage; F = Saturn V S-II stage; G = Saturn V S-IC stage; H = Spacecraft-Launch Adapter (SLA); I = aerodynamic shroud. Image credit: North American Aviation/NASA.

Near the end of its study, as it firmed up its spacecraft weight estimates, NAA determined that a single three-stage Saturn V, virtually identical to that used to launch Apollo lunar missions, could launch its Venus flyby spacecraft directly to Venus. The Saturn V S-IVB third stage would do the job of the S-IIB orbital launch stage. This led the company to reexamine its Mars flyby Earth-orbital launch scheme.

The company found that the heavier Mars flyby spacecraft could not launch directly onto its Mars flyby path if it were launched on a single three-stage Saturn V. It proposed instead that the Mars flyby spacecraft be split into two payloads — the CSM bearing the crew and the MC/PC combination — and that they be launched on a pair of three-stage Saturn Vs. CSM and MC/PC would then rendezvous and dock "in transit" soon after their S-IVBs placed them on course for Mars. 

The CSM would play the active role in the in-transit rendezvous. As Earth shrank behind it, its crew would separate the spacecraft from the S-IVB third stage that boosted it toward Mars, rendezvous and dock the MM/PC combination, and detach it from its S-IVB. After it moved a safe distance away, piloted flyby spacecraft instrument and antenna deployment and artificial-gravity spin-up could begin.

NAA provided a cost estimate for its 1973 Phase I Venus and 1975 Phase II Mars piloted flyby missions. The Venus mission would cost $2,301,700,000 between the 1 July 1967 contract go-ahead date and return to Earth on 19 December 1974. The Mars flyby without the Venus flyby would cost $3,439,500,000.

NAA representatives told MSC managers that its study had demonstrated that "existing and currently programmed hardware and facilities and systems contemplated for other NASA space flight programs can be used to achieve early Mars and/or Venus flyby missions." The company declared that "[f]ailure to take timely advantage of this opportunity could result in a delay in the achievement of advanced [orbital] and/or landing missions to Mars until the next century."


"One-Year Exploration Trip Earth-Mars-Venus-Earth," G. Crocco; paper presented at the 7th International Astronautical Federation Congress in Rome, Italy, 1-7 September 1956.

"Laboratory in Space," M. Yarymovych, NASA Headquarters; paper presented at the First Space Congress in Cocoa Beach, Florida, 20-22 April 1964.

"Future Effort to Stress Apollo Hardware," Aviation Week & Space Technology, 16 November 1964, pp. 48-51.

"An Evolutionary Program for Manned Interplanetary Exploration," M. W. Jack Bell; paper presented at the AIAA/AAS Stepping Stones to Mars Meeting in Baltimore, Maryland, 28-30 March 1966. 

Manned Mars and/or Venus Flyby Vehicles Systems Study Final Briefing Brochure, SID 65-761-6, North American Aviation, Inc., 18 June 1965.