CEV Credit - © Mark Wade

The Crew Exploration Vehicle (CEV) was NASA's planned manned spacecraft intended to carry human crews from Earth into space and back again from 2012 on.

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:

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From left to right, to same scale: CEV, Apollo CSM, Big Gemini, TKS, Soyuz, Shenzhou

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Parameter       Apollo CSM          CEV   Big Gemini           TKS         
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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
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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
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Lunar Landing:
SM Propellants   18,413 kg      9,300 kg 
Total Mass       30,629 kg     23,153 kg 
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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:

• The Block 1A would be used to rotate three to six crew members and cargo to the ISS and from the ISS. The spacecraft would require minimal modifications compared to the Block 2 baseline for this task. The Block 1A would have a total mass of 22,900 kg, including three crew, 400 kg of cargo, 8,300 kg of propellant, and 1,544 m/s of delta-V capability. However 6300 kg of the propellant was excess to the space station resupply mission. NASA vaguely says this unused propellant could be used for ISS reboost or in emergencies. In fact, this version could be flown on the ISS mission with a total mass as low as 16 metric tons, allowing it to be boosted on another booster (Atlas V, Delta 4, Ariane 5).
• The Block 1B would be as the Block 1A, but with all crew-related equipment removed and used unmanned to deliver 3,500 kg of pressurized cargo to the ISS and return an equivalent mass to Earth. The Block 1A would have a total mass of 22,900 kg, including 3,500 kg of cargo, 6,300 kg of propellant, and 1,098 m/s of delta-V capability. The same propellant caveats apply as the 1A. Eliminating unnecessary propellants would bring the launch mass down to 18,300 kg.
• The Cargo Delivery Vehicle (CDV), would replace the CM with an unpressurized cargo container. The SM would be fueled with the minimum propellants needed to maneuver the CDV to a rendezvous with the ISS. The ISS crew would then have to grapple the CDV with a remote manipulator arm and dock it to the station. The CDV was mainly a structural "strong-back" with a CBM for attachment to the ISS. Because the avionics for the other CEV variants were located within the CM, an avionics pallet would be required for the CDV. This pallet would support the avionics and provide the connection to the ATCS cooling system on the SM. The CDV was sized to transport two 1,500-kg unpressurized ORUs for the ISS. Examples of ORUs include Control Moment Gyroscopes (CMGs) and pump packages. The CDV would have a total mass of 19,112 kg, including the 12,200 kg CDV with 6,000 kg of net cargo. Two releasable cargo pallets on the CDV would be grappled by the stations manipulator arm and relocated to an ISS truss as required. Once the cargo had been removed from the CDV, it would undock from the station and perform an automated deorbit burn for burn-up and disposal in the ocean.
• The Block 2 was the baseline, following the profile noted above.
• The Block 3 was used as part of a reference Mars mission. It would transfer a crew of six between Earth and an undefined Mars Transit Vehicle at the beginning and end of the Mars exploration mission. The mission would require the CEV to be placed into a 55 km x 296 km initial orbit by the CEV, then maneuver for two days to a rendezvous and docking with the MTV in a circular orbit between 800 and 1200 km altitude. After the docking, the CEV would have to remain in quiescent storage for up to 30 months during the mission to Mars and back. 24 to 48 hours prior to Earth entry, the crew would enter the CEV and undock from the MTV. After undocking, the SM would be used to make a critical onboard-targeted, but ground-validated, burn to target for the proper entry corridor. The SM would then separate from the CM. Earth entry speeds for the CM from a nominal Mars return trajectory would be as high as 14 km/s, compared to 11 km/s for the Block 2 CEV. The changes and mass requirements for this version of the CEV compared to the Block 1 or 2 versions were left undefined as of the end of 2005. Obviously systems would have to modified for years of reliable storage on the way to Mars and back. Propellant requirements would be modest, and solar panels might not be needed.

Crew Size: 6. Orbital Storage: 180 days. Length: 9.84 m (32.28 ft). Basic Diameter: 5.50 m (18.00 ft). Maximum Diameter: 5.50 m (18.00 ft). Habitable Volume: 12.00 m3. Mass: 23,153 kg (51,043 lb). RCS Propellants: Lox/LCH4. Main Engine Thrust: 66.600 kN (14,972 lbf). Main Engine Propellants: LOX/LCH4. Main Engine Isp: 361 sec. Electrical System: Solar panels.

• CEV CM.  Other Designations: Crew Module. Class: Manned. Type: Spacecraft Module. Destination: Moon. Nation: USA. Agency: NASA.

  The Crew Module of NASA's Crew Exploration Vehicle was the re-entry capsule that also provided all of the planned living accommodation for up to six crew between earth and the ISS, the lunar surface, or an interplanetary Mars Transfer Vehicle. The spacecraft was to consist of a re-entry capsule capable of returning up to six crew at velocities of up to 14 km/s.

  The lunar mission version of the CEV CM, dubbed Block 2, was the baseline. It would transport four crew members from Earth to lunar orbit and return them to Earth. The CM would provide habitable volume for the crew, life support, docking and pressurized crew transfer to the Lunar Surface Access Module (LSAM), and atmospheric entry and landing capabilities. Upon return, a combination of parachutes and airbags provided for a nominal land touchdown with water flotation systems included for water landings following an aborted mission. Three main parachutes would slow the CM to a steady-state sink rate of 7.3 m/s. Just prior to touchdown, the ablative aft heat shield would be jettisoned and four Kevlar airbags deployed for a soft landing. After recovery, the CEV was to be refurbished and reflown with a lifetime up to 10 missions.

  A 41% scale-up of the Apollo Command Module shape with a base diameter of 5.5 m and sidewall angle of 32.5 degrees was selected for the outer moldline (OML). This configuration provided 29.4 m3 of pressurized volume and 12 to 15 m3 of habitable volume for the crew during transits between Earth and the Moon. The CEV CM operated at a nominal internal pressure of 0.648 atmospheres with 30 percent oxygen composition for lunar missions. However the pressure vessel structure was designed for a maximum pressure of one atmosphere. Operating at this higher pressure allowed the CEV to transport crew to the ISS without the use of an intermediate airlock. For the lunar missions, the CM would launch with a sea-level one atmospheric pressure, and the cabin was depressurized to 0.648 atmospheres prior to docking with the LSAM.

  The lunar CEV CM propulsion system provided vehicle attitude control for atmospheric entry following separation from the SM and range error corrections during the exoatmospheric portion of a lunar skip-entry return trajectory. A gaseous oxygen/ethanol bipropellant system was assumed with a total delta-V equivalent of 50 m/s.

  Other versions of the CM would fly before the Block 2 lunar version. The Block 1A CM would transport varying complements of crew and pressurized cargo to the International Space Station. The Block 1B CM would be unmanned, transporting pressurized cargo from earth to the ISS and back. The crew accommodations would be removed and replaced with secondary structure to support the cargo. The CEV would utilize a new androgynous Low-Impact Docking System (LIDS) to mate with other exploration elements and to the ISS. This would require changes to existing ISS equipment. It was proposed that two new docking adapters would replace the ISS Pressurized Mating Adapter (PMA) and Androgynous Peripheral Attachment System (APAS) adapters after Shuttle retirement.

  Finally, the Block 3 configuration would be used as a crewed transfer vehicle from the earth to a Mars Transfer Vehicle in Earth orbit. The crew would be six for the Block 3 CM, and this requirement drove the sizing of the design. Paradoxically, being some time in the future, no detailed design requirements were established by NASA for the Block 3 and detailed mass estimates were never derived.

  During ascent a Launch Abort System (LAS) would be used for emergencies, sized to pull the CM away from an errant launch vehicle with 10 G acceleration. The LAS concept was a tractor system that was mounted ahead of the CM. The exhaust nozzles were located near the top of the motor to reduce impingement loads on the CM. The LAS featured an active trajectory control system based on solid propellant, a solid rocket escape motor, forward recessed exhaust nozzles, and a CM adapter. The motor measured 76 cm in diameter and was 5.5 m in length. Eight canted thrusters minimized plume impingement on the CM. A star fuel grain minimized motor size and redundant igniters guaranteed system ignition in an emergency.

  The LAS provided abort from the launch pad and throughout powered flight of the booster first stage. The LAS would be jettisoned 20 to 30 seconds after second stage ignition. After the LAS was jettisoned, launch aborts for the crew would be accomplished using the SM propulsion system.

  The Block 1A CM would be used to rotate three to six crew members and cargo to the ISS and from the ISS. The Block 2 CM would require minimal modifications made to support ISS crew rotation. Initial mass for the baseline three-crew Block 1A CM variant would be 162 kg less than the Block 2 CM, with the differences as follows:

  • Remove EVA support equipment for one crew member (-3 kg);
  • Size galley, waste collection consumables, and soft stowage for 18 crew-days instead of 53 crew-days (-19 kg);
  • Removed one crew member and size personnel provisions for 18 crew-days (-238 kg);
  • Add ISS cargo (+400 kg);
  • Size oxygen, nitrogen, and potable water for 18 crew-days (-156 kg);
  • Sized RCS propellant for smaller vehicle mass and lower Delta-V (-145 kg)
  • Less growth allocation for lower vehicle dry mass (-4 kg).

  The Block 1B ISS Pressurized Cargo CM Variant would be used to deliver several metric tons of pressurized cargo to the ISS without a crew on board and return an equivalent mass of cargo to a safe Earth landing. Most components associated with providing crew accommodations would be removed and replaced with cargo. Initial mass for the uncrewed Block 1B CM variant would be 2,039 kg greater than the three-crew Block 1A crew rotation CM, with the assumed system modifications as follows:

  • Remove atmosphere contaminant (Carbon Dioxide (CO2), etc.) control equipment (-165 kg);
  • Remove EVA support equipment (-21 kg);
  • Remove galley, waste collection system, and cargo transfer bags (-84 kg);
  • Remove mass for personnel and personnel provisions (-580 kg);
  • Remove 500 kg of ISS cargo and ballast, and added 3,500 kg of ISS cargo (+3,000 kg);
  • Load oxygen, nitrogen, and water as needed for the pressurized cargo mission (-64 kg);
  • Increase RCS propellant for higher vehicle mass
  • Less growth allocation for lower vehicle dry mass (-54 kg).

The reference Mars mission utilized a Block 3 CM to transfer a crew of six between Earth and an MTV at the beginning and end of the Mars exploration mission. After a two-day flight leading to a docking with the MTV, the CEV would be configured to a quiescent state and remain docked to the MTV for the trip to Mars and back. Periodic systems health checks and monitoring was performed by the ground and flight crew throughout the mission. 24 to 48 hours prior to Earth entry after an 18 to 30 month mission, the crew would enter the CEV and undock from the MTV. After undocking, the CEV would make an onboard-targeted, but ground-validated, burn to target for the proper entry corridor. The CM would separate from the SM and then maneuvers to the proper attitude for a direct-guided entry to the landing site. Earth entry speeds from a nominal Mars return trajectory would be as high as 14 km/s, compared to 11 km/s for the Block 2 CEV. The CEV would perform a nominal landing at the primary land-based landing site and the crew and vehicle recovered.

Design Evolution of the NASA CEV CM

Using an improved blunt-body capsule for the CM was found to be the least costly, fastest, and safest approach for bringing ISS and lunar missions to reality. The key benefits for a blunt-body configuration were found to be lower weight, a more familiar aerodynamic design from human and robotic heritage (resulting in less design time and cost), acceptable ascent and entry ballistic abort load levels, crew seating orientation ideal for all loading events, and easier launch vehicle integration and entry controllability during off-nominal conditions.

The design and shape of the CM evolved in four design cycles throughout NASA's internal studies, beginning with an Apollo derivative configuration 5 m in diameter and a sidewall angle of 30 degrees. This configuration provided an OML volume of 36.5 m3 and a pressurized volume of 22.3 m3. The CM also included 5 g/cm2 of supplemental radiation protection on the cabin walls for the crew's protection. 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 the lunar surface direct landing mission scenario baseline at that point the CEV would be taken to the lunar surface.

A larger CEV was considered in Cycle 2 which grew the outer CM diameter to 5.5 m and reduced the sidewall angles to 25 degrees. Both of these changes substantially increased the internal volume. The pressurized volume increased by 75 percent to 39.0 m3 and the net habitable volume increased by over 50 percent to 19.4 m3. 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. Most of the system design parameters stayed the same for this cycle including the 5 g/cm2 of supplemental radiation protection.

Cycle 3 reduced the sidewall angles even further to 20 degrees in an effort to achieve monostability on Earth entry. The sidewall angle increased the volume further. Because the increases in volume were also increasing the vehicle mass, the height of the vehicle was reduced by 17 inches, reducing the height-to-width aspect ratio. This configuration showed the most promise in the quest for monostability, but the proper center of gravity was still not achieved. Analysis in this design cycle showed that the supplemental radiation protection could be reduced to 2 g/ cm2.

Cycle 4 was the final CEV design cycle and began after the decision was made to no longer consider the lunar surface direct mission. No supplemental radiation protection was included in the mass estimates for this design analysis due to results from a radiation study. The resulting Cycle 4 CM shape was a photographic scaling of the Apollo Command Module. The vehicle was 5.5 m in diameter and the CM had a sidewall angle of 32.5 degrees. The resulting CM pressurized volume was approximately 25 percent less than the Cycle 3 volume, but almost three times the internal volume as compared to the Apollo Command Module.

Although the baseline at the end of 2005 was a direct aerodynamic copy of the Apollo CM, NASA felt that further unspecified improvements to the Apollo shape could offer better operational attributes, especially by increasing the Lift-to-Drag (L/D) ratio, improving Center of Gravity (CG) placement, potentially creating a monostable configuration, and employing a lower angle of attack for reduced sidewall heating.

A CM measuring 5.5 m in diameter was chosen to support the layout of six crew without stacking the crewmembers above or below each other. A crew tasking analysis confirmed the feasibility of the selected vehicle volume. The available internal volume provided flexibility for future missions without the need for developing an expendable mission module. The vehicle scaling also considered the performance of the NASA-selected CLV, which was a four-segment SRB with a single SSME upper stage. The CM was scaled to maximize vehicle size while maintaining adequate performance margins on the CLV.

The choice of a primary land landing mode was primarily driven by a desire for land landing in the continental United States (CONUS) for ease and minimal cost of recovery, post-landing safety, and reusability of the spacecraft. However, the design of the CM would incorporate both a water- and land-landing capability. Ascent aborts or emergency returns from orbit would require the ability to land in water. In order for CEV entry trajectories from LEO and lunar return to use the same landing sites, it would be necessary for NASA to utilize skip-entry guidance on the lunar return trajectories. This technique was perfected by the Russians for their abortive manned lunar program in the 1960's. The skip-entry lunar return technique provided an approach for returning crew to a single CONUS landing site anytime during a lunar month. The Apollo-style direct-entry technique requires water or land recovery over a wide range of latitudes. The skip-entry included an exoatmospheric correction maneuver at the apogee of the skip maneuver to remove dispersions accumulated during the skip maneuver. The flight profile could be standardized for all lunar return entry flights. Standardizing the entry flights permitted targeting the same range-to landing site trajectory for all return scenarios so that the crew and vehicle experience the same heating and loads during each flight. This did not include SM disposal considerations, which must be assessed on a case-by-case basis.

Crew Size: 6. Orbital Storage: 180 days. Typical orbit: From ISS to lunar orbit. Length: 3.62 m (11.87 ft). Basic Diameter: 5.50 m (18.00 ft). Maximum Diameter: 5.50 m (18.00 ft). Habitable Volume: 12.00 m3. Mass: 9,506 kg (20,957 lb). RCS Propellants: Gox/Alcohol. RCS Impulse: 4,658 kgf-sec. L/D Hypersonic: 0.30. Electrical System: Solar panels.