The terms "conjunction class" and "opposition class" refer to the position of Mars relative to Earth during the Mars expedition. In the former, Mars moves behind the Sun as seen from Earth (that is, it reaches conjunction) halfway through the expedition; in the latter, Mars is opposite the Sun in Earth's skies (that is, at opposition) at the expedition's halfway point.
Conjunction-class expeditions are typified by low-energy transfers to and from Mars, each lasting about six months, and by long stays at Mars—roughly 500 days. Total expedition duration thus totals about 1,000 days. The long stay gives Mars and Earth time to reach relative positions that make a minimum-energy transfer from Mars to Earth possible. Von Braun opted for a conjunction-class expedition in The Mars Project.
Figure 4—Conjunction-class Mars missions include a low-energy transfer from Earth to Mars, a long stay at Mars, and a low-energy transfer from Mars to Earth. 1 - Earth departure. 2 - Mars arrival. 3 - Mars departure. 4 - Earth arrival. (Manned Exploration Requirements and Considerations, Advanced Studies Office, Engineering and Development Directorate, NASA Manned Spacecraft Center, Houston, Texas, February 1971, p. 1-7.)
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Opposition-class Mars expeditions have one low-energy transfer and one high-energy transfer separated by a short stay at Mars—typically less than 30 days. Total duration is about 600 days. This was the approach Lewis used in its 1959-1961 study. In the 1960s, most Mars expedition plans were opposition class.
Figure 5—Opposition-class Mars missions offer a short Mars stay but require one high-energy transfer, so they demand more propellant than conjunction-class missions. 1 - Earth departure (low-energy transfer). 2 - Mars arrival. 3 - Mars departure (high-energy transfer). 4 - Earth arrival. (Manned Exploration Requirements and Considerations, Advanced Studies Office, Engineering and Development Directorate, NASA Manned Spacecraft Center, Houston, Texas, February 1971, p. 1-8.)
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Because they require more energy, opposition-class expeditions demand more propellant. All else being equal, a purely propulsive opposition-class Mars expedition can need more than 10 times as much propellant as a purely propulsive conjunction-class expedition. This adds up, of course, to a correspondingly greater spacecraft weight at Earth-orbit departure.
Therefore, the conjunction-class plan is attractive. However, the long mission duration is problematical, for it demands great endurance and reliability from both machines and astronauts, exposes any crew left in Mars orbit to risk from meteoroids and radiation for a longer period, and requires complex Mars surface and orbital science programs to enable productive use of the 500-day Mars stay.
Mars in California
NASA's Ames Research Center, a former NACA laboratory in Mountain View, California, also became involved in piloted Mars planning in the EMPIRE era. In 1963, Ames contracted with the TRW Space Technology Laboratory to perform a non-nuclear Mars landing expedition study emphasizing weight reduction. Robert Sohn supervised the study for TRW and presented the study's results at the 1964 Huntsville meeting. 26 Sohn's team targeted 1975 for the first piloted Mars landing.
TRW found that the biggest potential weight-saver was aerobraking. For its aerobraking calculations, it used the Rand Corporation's August 1962 "Conjectural Model III Mars Atmosphere" model, which posited a Martian atmosphere consisting of 98.1 percent nitrogen and 1.9 percent carbon dioxide at 10 percent of Earth sea-level pressure. This atmospheric density and composition dictated the spacecraft's proposed shape—a conical nose with dome-shaped tip, cylindrical center section, and skirt-shaped aft section. This shape was based on an Atlas missile nose cone. The TRW team's two-stage, 12.5-ton MEM would also use the nose-cone shape. All else being equal, a version of TRW's spacecraft for the 1975 Mars launch opportunity that used braking rockets at Mars and Earth would weigh 3,575 tons, while the company's aerobraking design would weigh only 715 tons.
Figure 6—TRW's 1964 Mars ship design, shaped like a missile warhead, sought to minimize required propellant by aerobraking in the Martian atmosphere. This cutaway shows the Mars lander and Earth Return Module inside the spacecraft. ("Summary of Manned Mars Mission Study," Robert Sohn, Proceeding of the Symposium on Manned Planetary Missions: 1963/1964 Status, NASA TM X-53049, Future Projects Office, NASA George C. Marshall Spaceflight Center, Huntsville, Alabama, June 12, 1964, p. 150.)
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TRW's Earth aerobraking system was the Earth Return Module, a slender half-cone lifting body carried inside the main spacecraft. A few days before Earth encounter the crew would enter the Earth Return Module and separate from the main spacecraft. The Earth Return Module would enter Earth's atmosphere as the main spacecraft flew past Earth into solar orbit.
The TRW study proposed a lightweight artificial gravity system—a 500-foot tether linking the main spacecraft to the expended booster stage that pushed it from Earth orbit—which would, it calculated, add less than 1 percent to overall spacecraft weight. The resultant assemblage would spin end over end to produce artificial gravity. TRW reported that NASA Langley had used computer modeling to confirm this design's long-term rotational stability. 27
TRW found that Earth-Mars trajectories designed to reduce spacecraft weight at Earth departure would result in high reentry speeds at Earth return. For example, an Earth Return Module would reenter at 66,500 feet per second at the end of a 1975 Mars voyage, while one returning after a 1980 mission would reenter at almost 70,000 feet per second. TRW found that available models for predicting atmospheric friction temperatures broke down at such speeds. 28 For comparison, maximum Apollo lunar-return speed was "only" 35,000 feet per second.
Reentry speed could be reduced by using rockets. TRW found, however, that including enough propellant to slow the entire spacecraft from 66,500 feet per second to 60,000 feet per second would boost spacecraft weight from 715 tons to 885 tons. Slowing only the Earth Return Module by the same amount would increase overall spacecraft weight to 805 tons.
The study proposed a new alternative—a Venus swingby at the cost of a modest increase in trip time. A ship returning from Mars in 1975 could, the study found, cut its Earth reentry speed to 46,000 feet per second by passing 3,300 kilometers over Venus's night side. A Venus swingby during flight to Mars in 1973 would allow the ship to gain speed without using propellant and thus arrive at Mars in time to take advantage of a slower Mars-Earth return trajectory. According to TRW's calculations, Venus swingby opportunities occurred at every Mars launch opportunity. 29
Building on Apollo
By the end of the June 1964 Marshall Mars symposium, early flyby detractor Maxime Faget had come to see some merit in the concept. In a panel discussion chaired by Heinz Koelle, Faget declared that "we should, I think, consider a flyby . . . if we undertake a flyby we really have to face the problems of man flying out to interplanetary distances . ...I think we have to undertake a program that will force the technology, otherwise we will not get [to Mars] in my lifetime . . . ." 30
Von Braun, also a panel member, added, that "I think [piloted] flyby missions, particularly flybys involving [automated] landing probes . . . would be invaluable .... One such flight, giving us more information on what to expect . . . on the surface of Mars, will be extremely valuable in helping us in laying out the equipment for the landing ... that would follow the first flyby flight." 31
Von Braun then implicitly announced an impending shift in NASA advanced planning. "I am also inclined to believe," he said, "that our first manned planetary flyby missions should be based on the Saturn V as the basic Earth-to-orbit carrier. The reason is that, once the production of this vehicle is established and a certain reliability record has been built up, this will be a vehicle that will be rather easy to get." Von Braun's statement acknowledged that a post-Saturn rocket appeared increasingly unlikely. 32 In an outline of future plans submitted to President Lyndon Johnson's Budget Bureau in late November 1964, NASA stated that the post-Saturn rocket should receive low funding priority, and called for post-Apollo piloted spaceflight to be focused on Earth-orbital operations using technology developed for the Apollo lunar landing. 33
The 1964 decision to use Apollo technology for missions after the lunar landing could be seen as a rejection of post-Apollo piloted Mars missions. Historian Edward Ezell wrote in 1979 that the "determinism to utilize Apollo equipment for the near future was very destructive to the dreams of those who wanted to send men to Mars." 34 As if to emphasize this, the amount of funding applied to piloted planetary mission studies took a nose dive after November 1964. In the 17 months preceding November 1964, $3.5 million was spent on 29 piloted planetary mission studies. Between November 1964 and May 1966, NASA contracted for only four such studies at a cost of $465,000. 35
Mars planners were not so easily discouraged, however. After EMPIRE, and concurrent with UMPIRE, a Marshall Future Projects Office team led by Ruppe commenced an in-house study to look at using Apollo hardware for Mars exploration. Ruppe's study report, published in February 1965, found that piloted Mars flyby missions would be technically feasible in the mid- to late-1970s using Saturn rockets and other Apollo hardware. 36 The report's flyby spacecraft design used hardware already available or in an advanced state of development. Two RL-10 engines would provide rendezvous and docking propulsion, for example, and an Apollo Lunar Module descent engine would perform course corrections.
A pressurized hangar would protect a modified Apollo CSM during the interplanetary voyage. The hangar would also provide a shirt-sleeve environment so that the astronauts could act as in-flight caretakers for five tons of automated probes, including "landers, atmospheric floaters, skippers, orbiters, and possibly probes . . . to perform aerodynamic entry tests [of] designs and materials." 37 The last of these would, Ruppe wrote, provide data to help engineers design the piloted Mars landers to follow. His report drew on the UMPIRE conclusions when it stated that
significant reduction of initial mass in Earth orbit is possible if we can use aerodynamic braking at Mars or refueling there, but these methods assume a knowledge about . . . the Martian atmosphere, or about Mars surface resources which just is not available. The first venture, still assuming that we are not very knowledgeable . . . would probably transport 2 or 3 men to the surface of Mars for a few days ... [at a cost of] a billion dollars per man-day on Mars. If the physical properties of Mars were well known, we could think . . . of the first landing as a long-duration base, reducing cost to less than 10 million dollars per man-day. 38
The three-person flyby crew would live in a spherical habitat containing a radiation shelter and a small centrifuge for maintaining crew health (the study rejected artificial gravity systems that rotated the entire craft as being too complex and heavy). Twin radioisotope power units on extendible booms would provide electricity.
The mission would require six Saturn V launches and one Saturn IB launch. Saturn V rocket 1 would launch the unpiloted flyby spacecraft; then Saturn V rockets 2 through 5 would launch liquid oxygen tankers. The sixth Saturn V would then launch the Earth-departure booster, a modified Saturn V second stage called the S-IIB, which would reach orbit with a full load of 80 tons of liquid hydrogen but with an empty liquid oxygen tank. Ruppe wrote that solar heating would cause the liquid hydrogen to turn to gas and escape; to ensure that enough remained to boost the flyby craft toward Mars, the S-IIB would have to be used within 72 hours of launch from Earth.