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CEV CM

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CEV Credit: © Mark Wade

AKA: Crew Module. Status: Study 2006. Gross mass: 9,506 kg (20,957 lb). Height: 3.62 m (11.87 ft). Diameter: 5.50 m (18.00 ft).

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. Habitable Volume: 12.00 m3. RCS total impulse: 4,658 kgf-sec.