When Mike Griffin was appointed NASA Administrator, he threw out the previous elaborate plans for evolutionary development of a Crew Exploration Vehicle through a long, expensive, 'spiral' development process. Instead he obtained White House backing to plunge ahead using existing technology and NASA's best judgment. The imaginative proposals from industry were largely ignored, except where they supported NASA's own conclusions. Lockheed and Northrop were notified that they had 'won' the CEV design competition in June 2005, except that they would be only be allowed to make final bids based on the design dictated by NASA. NASA's own configuration was called 'Apollo on steroids'. Block 1 versions of the CEV would be used initially to provide access to the International Space Station after the retirement of the Space Shuttle in 2011. Thereafter it would provide the earth return vehicle for missions to the moon (Block 2, by 2020) or Mars (Block 3, by 2030+?).
NASA proposed a spacecraft with 23% less total mass than the Apollo CSM, 25 metric tons, but with a greater basic diameter of 5.5 m, rather than Apollo's 3.9 m. The re-entry vehicle was a 41% scaled up version of the Apollo command module. This would have over three times the internal volume and double the surface area of the Apollo capsule, but NASA claimed its mass could be limited to only 64% more than the Apollo design. Despite the increase in volume and mass, it would provide accommodation for only four to six crew (versus three to five in Apollo), plus up to 400 kg of payload that could be returned from orbit. An unmanned version of the capsule, with all crew provisions removed, could deliver or return up to 3500 kg of cargo to the International Space Station.
The service module was stubbier and lighter than the Apollo CSM, and powered by a 6800 kgf liquid oxygen/methane engine. The same propellant combination would be used in the reaction control system. The ascent stage of the planned Lunar Surface Access Module would use the same rocket engines. Liquid oxygen/methane had only been investigated seriously by the Russians in the 1990's, so this decision represented the only really new technology in NASA's CEV design and the major risk in its development schedule. The choice was driven by NASA plans to generate methane from the Martian atmosphere on future manned expeditions. The mass of the service module was kept down by the higher specific impulse of the liquid oxygen/methane propellants compared to the storable propellants used in the original Apollo. More importantly, for the lunar landing scenario, the CEV would be required to make only the Trans-Earth injection maneuver to bring the crew home. In the Apollo scenario, the CSM had to brake both the CSM and lunar module into lunar orbit, as well as make Trans-earth injection for the CSM.
The CEV would be launched into earth orbit by the Crew Launch Vehicle, a shuttle-derived two-stage rocket consisting of a single Shuttle RSRM solid booster as the first stage and a new second stage, 5.5 m in diameter, using Lox/LH2 propellants and powered by a single SSME.
The CEV, while breaking no radical new technological ground, did seem ambitious and could be a weight-control challenge. NASA seems to be looking to duplicate the performance of the Apollo CSM at an overall mass 20% less, while providing more space for the crew and generous provisions for cargo taken to orbit and back in the return capsule. Apparently the allowance for weight growth was only 2.4%, while NASA's own experience on manned spacecraft had been over 20%. A comparison with comparable manned spacecraft shows the differences and challenges:
Parameter Apollo CSM CEV Big Gemini TKS -------------------------------------------------------------------- CM Mass 5,806 kg 9,506 kg 5,227 kg 3,800 kg SM Dry mass 6,240 kg 4,380 kg 9,000 kg 4,910 kg Total Mass Dry 12,046 kg 13,886 kg 9,000 kg 8,710 kg Engine thrust 9,979 kgf 6,800 kgf 6,800 kgf 800 kgf Engine Isp 314 sec 362 sec 273 sec 291 sec Spacecraft dV 2,800 m/s 1,742 m/s 350 m/s 700 m/s Crew 3 to 5 4 to 6 9 3 -------------------------------------------------------------------- Station Resupply: Payload up to 2,000 kg* 3,500 kg 2,500 kg 4,528 kg Propellant up to 2,700 kg 6,900 kg 1,900 kg 3,822 kg Total Mass 14,700 kg 22,900 kg 15,950 kg 17,510 kg ----------------------------------------------- Lunar Landing: SM Propellants 18,413 kg 9,300 kg Total Mass 30,629 kg 23,153 kg -----------------------------------------------
On June 13, 2005, NASA announced the down-select of two contractors: Lockheed-Martin and the team of Northrop-Grumman and Boeing. A single contractor would be selected without prototyping or flight-testing in 2006, so that the spacecraft could be available by 2010 as a shuttle replacement. Initial indications were that crew size would be increased to six, and launch mass to 30 metric tons, signaling the new Administrator's stated preference to develop a Shuttle-derived vehicle as the CEV booster, and not try to man-rate existing Delta IV or Atlas V EELV's. The CEV was downgraded to under 25 metric tons and four crew by September 2005, putting it within the capability of heavy-lift versions of the EELV's.
By the end of 2005 it was clear that NASA was going to dictate the design of both the CEV and its shuttle-derived launch vehicle. It looked like the errors of the original Apollo program would be repeated. These included a decision process that proceeded from false assumptions; and not taking the minimum-mass approach.
When NASA was instructed to go to the moon in April 1961, the Apollo capsule design concept had already been selected. It was to be a general-purpose spacecraft, with a crew of three. The crew of three was selected for no other reason than NASA assumed that one crewmember would have to be awake at all times to tend unreliable automatic spacecraft systems, ergo three eight-hour shifts. The capsule had to have more space than required just to house three astronauts - enough for the crew to don and doff spacesuits on missions of up to two weeks. NASA's approach to the lunar landing mission took as its first assumption that the return capsule would be the Apollo, with a mass of five metric tons for three men.
However, the mission as it was framed was to 'land a man safely on the moon'. Kennedy did not say how many. Rather than lose a few months already expended on defining the design of the three-man Apollo capsule, NASA selected that as the baseline. After over a year of torturous internal argument, they decided to use a complex lunar orbit rendezvous mission profile, which meant two men would be landed on the moon. But if two men on the moon was the objective, the return capsule could be correspondingly reduced in size, with proportionate reductions in the size of the overall spacecraft, the rocket stages, and the launch vehicle. Using the same Saturn V launch vehicle and a two-man capsule, a direct mission to the moon could have been made, eliminating the lunar module and the complex and risky lunar-orbit rendezvous mission mode.
However there were further methods of weight reduction. General Electric proposed an alternative Apollo design that cut the mass of a three-man return vehicle to three metric tons. This put all systems and space not necessary for re-entry and recovery outside of the re-entry vehicle, into a separate jettisonable 'mission module', joined to the re-entry vehicle by a hatch. Every gram saved in this way saved two or more grams in overall spacecraft mass - for it was mass that did not need to be protected by heat shields, supported by parachutes, or braked on landing. Weight was also saved through use of a re-entry vehicle of the highest possible volumetric efficiency (internal volume divided by hull area). Theoretically this would be a sphere. But re-entry from lunar distances required that the capsule be able to bank a little, to generate lift and 'fly' a bit. This was needed to reduce the G forces on the crew to tolerable levels. Such a maneuver was impossible with a spherical capsule. After considerable study, the optimum shape was found to be the Soyuz 'headlight' shape - a hemispherical forward area joined by a barely angled cone (7 degrees) to a classic spherical section heat shield.
This design concept meant splitting the living area into two modules - the re-entry vehicle, with just enough space, equipment, and supplies to sustain the crew during re-entry; and a mission module. As a bonus the mission module provided an airlock for exit into space and a mounting area for rendezvous electronics.
The end result of this design approach was remarkable. The Apollo capsule designed by NASA had a mass of 5,000 kg and provided the crew with six cubic meters of living space. A service module, providing propulsion, electricity, radio, and other equipment would add at least 1,800 kg to this mass for the circumlunar mission. The General Electric spacecraft provided the same crew with 9 cubic meters of living space, an airlock, and the service module for the mass of the Apollo capsule alone!
The modular concept was also inherently adaptable. By changing the fuel load in the service module, and the type of equipment in the mission module, a wide variety of missions could be performed. The same concept was adopted by the Soviets for their Soyuz capsule, and the Chinese for their Shenzhou. The superiority of this approach was clear to see: the Soyuz remained in use 40 years later, while the Apollo was quickly abandoned.
Incredibly, NASA made the same mistakes again, forty years later. But this time in spades. First, an assumption was made that a six-person crew would be needed to be returned from a Mars mission. Everything else derived from this initial assumption. Now in the first place, there was no logical basis for selecting six. Dozens of Mars expedition studies have been made over the years, with anywhere from three to seventy crew being recommended. Furthermore, there was no logic to insisting that the entire crew come back in a single re-entry vehicle. Finally, this Mars mission with six crew that have to come home together, was totally unfunded, unengineered, and at least twenty years in NASA's future. It was in 1968 as well. So using this ever-receding dream as the basis for your utility spacecraft design was jaw-dropping.
At this point in its internal studies, NASA was assuming a direct flight to the moon. A large booster would fire the CEV toward the moon, it would land directly on the surface, and then it would fire an ascent stage and return to earth. This was the least complex and the most operationally efficient - trips from the earth to the moon and back could be made at any time. Furthermore, the capsule would be a maximum of 5.0 m in diameter, so that it could be launched by the Delta IV or Atlas V EELV boosters just put into service by the United States. But then NASA claims it found that: "Layouts for a crew of six and the associated equipment and stowage were very constrained and left very little habitable volume for the crew. It was determined that the internal volume for the CM was too small, especially for a surface direct mission where the CEV would be taken to the lunar surface". So NASA moved on to Design Cycle 2, with a 5.5 m diameter, where: "The desire in this design cycle was to provide enough interior volume for the crew to be able to stand up in and don/doff lunar EVA suits for the surface direct mission".
So here we were right back at Apollo. A large crew defined for no particular reason. A requirement was conceived that the re-entry vehicle had to provide all of the space the crew would need for the entire mission. And finally, based only on "constrained" layouts, NASA decides Not Invented Here launch vehicles would be unsuitable and decides only its own special-design boosters would do.
Now NASA had a capsule designed for six crew at some future date, but it discovered that it would need to only fly three crew to the ISS and four to the moon and back. Did NASA junk the 5.5 m design, and declare that it could use existing boosters after all? No, it plowed ahead with this six-crew capsule, and jinned up launch vehicles that would require existing shuttle components, facilities, and staff to make them fly. So it was clear that the mission was not to get to the moon, or support the ISS, but only to preserve NASA and NASA contractor jobs and facilities.
This doesn't even touch on the matter on the innovative designs that were suggested in the first round of CEV proposals that NASA would not even comment on. The same approach was used in Apollo. First, proposals from industry were solicited. In both the Apollo and CEV cases these were imaginative, innovative, and incorporated all of the lessons of hundreds of millions of dollars of advanced research funded not just by NASA, but also by industry and the US Air Force. Superior contractor designs using the Soyuz-type separate orbital module or a winged spaceplane approach were made in both cases. In the case of the CEV, the team that designed and flew SpaceShipOne on the first civilian manned spaceflight offered to build a complete (four-crew) air-launched booster and spacecraft that would do the job for one-fortieth of the CEV/CLV cost!. In both the Apollo and CEV cases the contractors were thanked, and NASA then proceeded with its own in-house government design. This was then suitably tweaked until it would passed the Congressional pork test.
CEV Technical Description
An earth-orbit-rendezvous/lunar-orbit rendezvous lunar landing mission was used to define the baseline design criteria for the CEV. The CEV for this mission was designated Block 2. Block 1 CEV's would support the International Space Station and Block 3 CEV's would be used for return from a Mars mission.
The Block 2 baseline lunar scenario involved the placing in an earth parking orbit by a shuttle-derived heavy-lift Cargo Launch Vehicle of a trans-lunar injection (TLI) stage and an unmanned Lunar Surface Access Module (LSAM). Within 30 days, a Crew Launch Vehicle (CLV) would launch the CEV into orbit. The CEV would rendezvous and dock with the LSAM and TLI stage. The combined spacecraft would be put on a trans-lunar trajectory by the TLI stage, which would then be jettisoned. The LSAM would brake the combined CEV and LSAM spacecraft into lunar orbit. The four-man crew would transfer to the LSAM and descend to the surface, leaving the CEV in quiescent mode in lunar orbit for up to six months. NASA provided the CEV with enough extra maneuvering capability to make substantial lunar orbit plane changes (evidently about 300 m/s of delta-V was allocated for this purpose). This would be used to bring the CEV's orbit over the landing site in case an emergency return to earth was needed. Without this capability, the crew could return to earth only twice a month unless the landing site was near the equator or poles.
When the crew was ready to leave the moon, they would fire the LSAM ascent stage, and then rendezvous and dock with the unmanned CEV. The LSAM would be cast away, and the CEV would perform the trans-earth injection maneuver, its only major engine burn. On approach to earth, the CM would separate from the SM after being set up for the precise angle and position needed for re-entry at 11 km/s. The CEV would land in the continental United States, and to do this it would be necessary for NASA to utilize a skip-entry technique perfected by the Russians for their abortive manned lunar program in the 1960's. The CM would descend under parachutes, with final landing on hard ground being cushioned by air bags.
The following versions of the CEV were planned:
Crew Size: 6. Orbital Storage: 180 days. Habitable Volume: 12.00 m3.
AKA: Crew Exploration Vehicle.
Gross mass: 23,153 kg (51,043 lb).
Height: 9.84 m (32.28 ft).
Diameter: 5.50 m (18.00 ft).
Thrust: 66.60 kN (14,972 lbf).
Specific impulse: 361 s.