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).

Liftoff of Gemini V, 21 August 1965. Image credit: NASA
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 on 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. 

Star-Raker (1978)

Star-Raker (right), a single-stage-to-orbit space plane, parks next to a 747 at a conventional airport. Image credit: M. Alvarez/Rockwell International.

Elsewhere in this blog, I have described the 1970s joint NASA/Department of Energy Solar Power Satellite (SPS) studies (see "More Information" below). Had even a single SPS been assembled, it would have been by far the largest human construction project in space; it would have weighed more than 100 times as much as the 420-metric-ton (460-U.S.-ton) International Space Station. The SPS studies envisioned assembly of two such satellites per year between 2000 and 2030, bringing the total number in the SPS constellation to sixty. 

NASA envisioned boosting SPS components to low-Earth orbit (LEO) in the payload bays of massive reusable launch vehicles. One such launcher, Boeing's winged, two-stage Space Freighter, would have weighed about 11,000 metric tons (12,125 U.S. tons) at liftoff and delivered about 420 metric tons (463 U.S. tons) to LEO. For comparison, the two-stage Saturn V rocket used to place 77-metric-ton (85-U.S.-ton) Skylab into LEO weighed about 2800 metric tons (3086 U.S. tons) at liftoff.

The Space Freighter would have risen vertically from a launch pad and pointed itself generally toward the east. As its first stage, the Booster, expended its propellants, it would have separated. The second stage, the Orbiter, would then have ignited its engines to complete its climb to LEO. In orbit, it would have maneuvered to rendezvous and dock with a large space station designed specifically for handling SPS cargo modules.

The Space Freighter Booster would have been a fully reusable winged vehicle closely resembling the Space Freighter Orbiter. After Space Freighter Orbiter separation, the Space Freighter Booster would have turned, deployed jet engines, and flown to a long, wide runway at its launch site. 

To begin return to Earth, the Space Freighter Orbiter in LEO would have separated from the cargo-handling space station, then would have turned its tail forward and ignited rocket motors to slow down, lowering its orbit so that it intersected Earth's atmosphere. Following a fiery reentry, it would have landed on the runway near its launch pad. 

After launch pad, Orbiter, and Booster refurbishment, the two Space Freighter stages would have been hoisted vertical. After the Orbiter was placed atop the Booster's nose, a cargo module would have been loaded into its payload bay. The Space Freighter would then have been moved to a launch pad to begin another flight. Launching parts for two SPS into LEO in a year would have required about 240 Space Freighter launches, or about one launch every 36 hours.

In October 1977, a team of 14 Rockwell International engineers studied a Space Freighter alternative. The Star-Raker space plane, 103 meters (310 feet) long with a wing span of about 93 meters (280 feet), would have carried a maximum of 89.2 metric tons (98.3 U.S. tons) of cargo into LEO. More than 1100 flights would have been required each year to support the SPS program, or about one launch every eight hours.

In its fully developed form, however, Star-Raker would have had important advantages over Space Freighter which might have made its required high flight rate feasible. For example, it would have begun its flights to LEO by taking off horizontally from a conventional 2670-to-4670-meter-long (8000-to-14,000-foot-long) runway at virtually any civilian or military airport capable of supporting 747 or C-5A Galaxy cargo planes. No specialized launch and landing site would have been required.

Every bit as important, Star-Raker would have been capable of flying routinely between such airports. The Rockwell team explained that this would "reduce the number of operations required to transport material and equipment from their place of manufacture on Earth to [LEO]." For example, rolls of solar cell blankets would not need to be shipped by train, barge, or plane to a specialized launch and landing site; they would, potentially, need only be transported to a local airport for Star-Raker pickup.

Though the 1977-1978 Star-Raker study focused on its possible use in the Department of Energy/NASA Solar Power Satellite program, Star-Raker would have had potential as a general-purpose space cargo plane. In the image above, three Star-Rakers, their nose sections hinged back to expose their cargo bays, take on payloads bound for destinations ranging from low-Earth orbit to deep space. Image credit: M. Alvarez/Rockwell International.

David Reed, an engineer at North American Rockwell (NAR), as the company was then known, originated the Star-Raker concept in 1968, as NASA began earnest efforts to develop a reusable Space Shuttle. Key elements of the concept had been proposed — and rejected — earlier in the 1960s decade. These included wings packed with lightweight structurally integral tanks holding liquid hydrogen fuel and liquid oxygen oxidizer and a complex jet engine/rocket engine propulsion system.

The 1968-1969 study determined that, as it burned the propellants in its wings and maneuvered through ascent from subsonic speed to Mach 6 (six times the speed of sound), aerodynamic pressure on its structure would become excessive. This led NAR to examine wing designs developed in 1970 for the proposed (and subsequently abandoned) U.S. Supersonic Transport program. 

A "tridelta flying wing" design appeared to solve the pressure problem; by then, however, NASA had narrowed its Shuttle design requirements, excluding Star-Raker from consideration. NAR continued Shuttle studies and became Shuttle prime contractor in July 1972. 

Rockwell revived study of the tridelta flying wing Star-Raker as SPS studies ramped up in 1976. The Star-Raker study that began in October 1977, led by Reed and performed for NASA Marshall Space Flight Center (MSFC) in Huntsville, Alabama, continued into late 1978, yielding the design described in this post.  

The 1977-1978 study benefited from computer modeling that enabled Rockwell to further refine Star-Raker wing shape and flight profile. It also allowed Reed's team to take more fully into account the benefits of propellant-saving "lifting ascent." 

Star-Raker's propellants, liquid hydrogen and liquid oxygen, were not typically found at airports in 1977-1978; this remains true in 2020. The Star-Raker study team might have assumed that airports would evolve to provide them by the time SPS cargo flights began in 2000. This would, perhaps, not have been an unreasonable assumption, given that the 30-year SPS program was expected to create a lucrative new industry spanning the continental United States. 

One Star-Raker takes off as another undergoes airport servicing. With its landing gear extended, Star-Raker ground clearance would have been 1.52 meters (five feet). Image credit: M. Alvarez/Rockwell International.

For the 1977-1978 study, however, they hedged their bets by assuming that liquid hydrogen fuel would be available at airports only in sufficient quantities for airport-to-airport subsonic air-breathing jet engine Star-Raker flights. Liquid oxygen would, of course, not have been required. Flights to LEO, which would have needed both propellants in large quantities, would have begun on a runway at NASA's Kennedy Space Center (KSC) in Florida, at Vandenberg Air Force Base, California, or at any other launch sites the U.S. might have deigned to establish. 

The propellant tanks in Star-Raker's wings would have been approximately conical in shape. They would have extended from the space plane's body to its wing tips and been designed to strengthen the wings with minimal weight penalty. They would have been reinforced with regularly spaced "cell web" walls. Foam-filled glass-fiber honeycomb would have surrounded the tanks, defining Star-Raker's shape.

The Rockwell team described in detail a Star-Raker flight from KSC to 556-kilometer-high (345-mile-high) LEO and back to a U.S. airport. It would have begun with arrival at KSC of a Star-Raker space plane loaded with cargo bound for LEO at the end of a subsonic flight from a conventional airport. 

Following a limited airplane-type checkout, crews would have installed three sets of jettisonable orbital-takeoff main landing gear, each with eight wheels, and pumped liquid hydrogen and liquid oxygen propellants into Star-Raker's tanks. Fully loaded with propellants and cargo and with its orbital-takeoff gear attached, Star-Raker would have weighed about 1935 metric tons (2130 U.S. tons). 

Star-Raker would have lifted off from the runway at a speed of 415 kilometers per hour (260 miles per hour) under "supercharged afterburner" power from its 10 multicycle jet engines. The Rockwell team explained that it had consulted with leading jet engine manufacturers to arrive at its jet engine design; these included General Electric, Pratt & Whitney, Aerojet, Marquardt, and Rocketdyne. The resulting engine was more a wish list than a firm design, though it was an informed wish list. 

The Rockwell team envisioned four operational cycles for its jet engine ranging from conventional turbofan to ramjet. Liquid hydrogen would have been used to cool the engine and then burned as fuel. Large, slot-shaped inlets on the underside of Star-Raker's wings, arranged in two groups of five on either side of the space plane's body, would have funneled air to the engines, which would have been mounted at the wing trailing edge. The inlets would have been equipped with "ramp" doors that could close partially or fully to moderate or halt airflow.

Shortly after leaving the ground, the space plane's crew would have dropped the three sets of orbital-takeoff landing gear (they would have lowered to the ground on parachutes for recovery and reuse), then would have retracted its nose and main landing gear. The space plane would then have switched its jet engines to turbofan power, climbed to 6100-meter (20,000-foot) cruise altitude, and increased its speed to Mach 0.85. It would have turned due south and, over the next hour and fifty minutes, flown directly to Earth's equator.

Star-Raker would have flown to the equator and turned east so that it could get a boost from Earth's rotational velocity, which at our planet's midriff can, in theory, add about 1600 kilometers (1000 miles) per hour to the orbital velocity of ascending launch vehicles. 

In addition, and more importantly, the turbofan flight to the equator would have amounted to a plane-change maneuver; that is, it would have enabled Star-Raker to reach equatorial LEO without performing the rocket-propelled plane-change maneuver in LEO required if Star-Raker flew directly to orbit from a non-equatorial launch site, such as KSC. The Rockwell team hoped that this would save propellants, enabling an increase in cargo weight.

Following the eastward turn, the space plane would have climbed to 13,710 meters (45,000 feet) under supercharged afterburner power, then would have begun a shallow dive to 11,280 meters (37,000 feet). During the powered dive, a propellant-saving maneuver, Earth's gravity would have helped it to break the sound barrier and accelerate to Mach 1.2. 

Go for orbit: the Star-Raker space plane design included 10 multicycle air-breathing jet engines, three high-pressure rocket engines akin to the Space Shuttle Main Engine, and two advanced Orbital Maneuvering System rocket engines. In the image above, the 10 jet engines are throttling up to begin the transition to supersonic flight. Image credit: M. Alvarez/Rockwell International.

Star-Raker would then have begun ascent to orbit in earnest, with a supersonic climb to 29 kilometers (18 miles). During this phase, the space plane's jet engines would have throttled up to "full ramjet" power, accelerating it to Mach 6.2. Throughout its climb to orbit, Star-Raker would have maneuvered to put to good use lift provided by its wings. 

Upon reaching Mach 6.2, the three rocket motors in Star-Raker's tail would have ignited, adding rocket power to ramjet power. The three engines, with a combined thrust of 1.45 million kilograms (3.2 million pounds), would have drawn liquid hydrogen from a sturdy tank located at the aft end of the long, narrow Star-Raker cargo bay. The tank, to which the engines would have been mounted, would have served as the load path that would have distributed their thrust to the space plane's structure.

At Mach 7.2, Star-Raker would have switched to full rocket power. As it throttled up the rocket motors to full thrust, it would have shut down the jet engines and closed completely their air inlet doors. 

When Star-Raker reached a 51-kilometer-by-556-kilometer (32-mile-by-345-mile) equatorial orbit, the main rocket motors would have shut down. At apogee, the high point in its orbit, the crew would have ignited the twin advanced Orbital Maneuvering System (OMS) engines at the base of its tail to raise its perigee (orbit low point) and circularize its orbit. Upon attainment of circular equatorial orbit, Star-Raker would have used the OMS to maneuver to a rendezvous with the SPS cargo-handling space station.

Star-Raker in low-Earth orbit. Image credit: M. Alvarez/Rockwell International.

The weight of cargo Star-Raker could carry would depend on its mission profile. For the profile described here, cargo weight delivered to orbit would have amounted to only about 48.6 metric tons (53.6 U.S. tons). The aerodynamic flight to the equator under jet power, meant to steal some of the Earth's rotational energy and avoid a plane change maneuver in LEO, had under close examination turned out to be expensive. 

The Rockwell team proposed improving the equatorial profile's payload performance by loading liquid oxygen at the equator, either during flight using a new-design tanker aircraft, or after a landing at an equatorial facility with an adequate runway, orbital-takeoff gear attachment and recovery capability, and ability to provide liquid oxygen. Either approach would, however, have complicated Star-Raker operations.

To unload cargo, Star-Raker would have swung its nose, which would have contained its crew compartment, sideways out of the way, exposing one end of its six-meter-high-by-six-meter-wide-by-43-meter-long (20-foot-high-by-20-foot-wide-by-141.5-foot-long) cargo bay. The bay's arched ceiling would have made it a point of structural strength, not weakness, in the Star-Raker design.

The crew would have moved to the rear of the crew compartment to assist with cargo transfer. Windows at the rear of the two-deck crew compartment would have provided a 121° field of visibility. 

The Rockwell team did not describe its cargo transfer system in any detail, though it is clear that Star-Raker would not have docked in the conventional sense. Brief mention was made of a transfer rail system in the cargo bay that would have linked to equivalent rails on the space station.

Return to Earth would have begun with cargo bay closure. After moving away from the space station, the crew would have turned Star-Raker so that its tail faced in its direction of orbital motion, then would have fired its OMS engines to slow down. 

Maximum deceleration during the unhurried shallow-angle reentry would have reached no more than 2.3 gravities. Star-Raker would, in general, have experienced reentry temperatures lower than the Space Shuttle Orbiter, though nose and wing leading-edge temperatures were expected be somewhat higher. The higher leading-edge temperature was attributable to its relatively blunt shape. 

The Rockwell team proposed two types of reusable Thermal Protection System (TPS) for Star-Raker. Both would have been mounted on an outer facing sheet covering a honeycomb layer. The honeycomb layer would in turn have been attached to an inner facing sheet covering the honeycomb core that surrounded the propellant tanks.

The first TPS design closely resembled that baselined for the Space Shuttle Orbiter. Ceramic tiles individually molded and milled to match Star-Raker's curves would have been glued to fabric strain-isolator pads affixed to the outer facing sheet. 

The second TPS design, similar to one developed for the B-1 Bomber, was more complex. Metal panels — titanium-aluminum for low-temperature areas and "superalloy" for high-temperature areas — would have been attached to the outer facing sheet using flexible standoffs. The standoffs would have permitted the overlapping panel edges to slide over each other as they grew hot and expanded or cooled and contracted. Foil-wrapped thermal insulation blankets affixed to the outer facing sheet would have provided additional thermal protection.

Both TPS designs would have included a system for detecting breaches in the TPS. The Rockwell team provided no details of its design and did not describe what the crew might do if a breach were detected.

Star-Raker on approach. Image credit: M. Alvarez/Rockwell International.

When Star-Raker slowed to Mach 6, it would have begun cross-range maneuvers designed to shed energy and slow it to Mach 0.85. The crew would then have opened the inlet ramps and started "some" of its jet engines. 

The Rockwell team provided the space plane with enough liquid hydrogen to permit a 556-kilometer (345-mile) subsonic cruise and two powered landing attempts. Landing velocity would have been about 215 kilometers per hour (135 miles per hour). At wheels stop at an airport capable of supporting a cargo 747 or a C-5A Galaxy, Star-Raker would have weighed about 281 metric tons (310 U.S. tons).

Star-Raker weights given in this flight description are based on data the Rockwell team generated in the period spanning December 1977-January 1978. In February-March 1978, NASA MSFC and NASA Langley Research Center (LaRC) in Hampton, Virginia, reviewed the Rockwell team's Star-Raker weight numbers. 

The NASA centers found that Rockwell's estimates were low if "normal" technology were assumed and high if "acceleration" (advanced) technology were assumed. Whereas Rockwell had placed Star-Raker's "dry" weight with orbital-takeoff gear at 293.5 metric tons (323.5 U.S. tons), MSFC/LaRC determined that, with normal technology and a 10% cushion for weight growth during development, Star-Raker would weigh 407.6 metric tons (449.3 U.S. tons) without propellants; with advanced technology and the cushion, it would weigh only 257.6 metric tons (284 U.S. tons). 

The Rockwell team and NASA MSFC engineers met in May 1978 to try to reconcile the weight estimates. They made one important change in Star-Raker's flight profile: they abandoned the subsonic flight to the equator in favor of a KSC launch and direct climb to a 556-kilometer (345-mile) LEO inclined 28.5° relative to Earth's equator (that is, the latitude of KSC). 

The NASA and Rockwell teams settled on a Star-Raker weight without propellants (but with orbital-takeoff gear and 10% cushion) of 330.4 metric tons (364.2 U.S. tons). As it began ascent to orbit on a KSC runway, the space plane would have weighed 2280.5 metric tons (2514 U.S. tons). Of this, Star-Raker's maximum weight, 89.2 metric tons (98.3 U.S. tons) would have comprised cargo for the SPS project.


Independent Research and Development Data Sheet, Project Title: Earth-to-LEO Transportation System for SPS, Rockwell International, 15 December 1978.

"Star-Raker: An Airbreather/Rocket-Powered, Horizontal Takeoff Tridelta Flying Wing, Single-Stage-to-Orbit Transportation System," SSD 79-0082, D. Reed, H. Ikawa, and J. Sadunas, North American Rockwell Space Systems Division; paper presented at the American Institute of Aeronautics & Astronautics Conference on Advanced Technology for Future Space Systems in Hampton, Virginia, 8-11 May 1979.

More Information

Electricity from Space: The 1970s DOE/NASA Solar Power Satellite Studies

NASA Johnson Space Center's Shuttle II (1988)