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Part of Orion CEV
Credit: © Mark Wade
American manned spacecraft module. Study 2006. The Service Module of NASA's Crew Exploration Vehicle provided basic consumables, control systems, and sufficient delta-V for return of the CEV from lunar orbit to the earth.

AKA: Service Module. Status: Study 2006. Thrust: 66.60 kN (14,972 lbf). Gross mass: 13,647 kg (30,086 lb). Unfuelled mass: 4,380 kg (9,650 lb). Specific impulse: 362 s. Height: 6.22 m (20.40 ft). Diameter: 5.50 m (18.00 ft).

The unpressurized SM provided propulsion, power, and other supporting capabilities to meet the CEV's in-space mission needs. The baseline CEV, dubbed Block 2, was sized for lunar missions carrying a crew of four. Block IA and IB versions of the CEV would transfer and return crew and cargo to the ISS and stay for 6 months in a quiescent state for emergency crew return. The lunar and ISS configurations shared the same SM, but the ISS mission had much lower Delta-V requirements (NASA used 1724 m/s as the required delta-V to return from lunar orbit to the earth, while simply maneuvering to and returning from the ISS would require only 350 m/s). Therefore the SM propellant tanks could be loaded with more propellant than needed for ISS missions. The additional propellants would provide flexibility in launch aborts and on-orbit phasing. Excess propellants could be used for ISS re-boost to adjust its orbit for decay. The SM could also be used as a tug to deliver only cargo to the ISS. In that case propellant would be offloaded, allowing up to six metric tons of payload to be delivered to the station.

A relatively undefined Block 3 CEV would deliver crews from the earth to an orbiting Mars Transfer Vehicle. In this case the SM would be required to be active for two days between launch and docking with the MTV, then quiescent for up to three years, then active again for a few hours in the critical earth re-entry phase.

The baseline Block 2 lunar CEV SM provided major translational maneuvering capability, power generation, and heat rejection for the CEV CM. The SM would be of stringer/ring frame construction with graphite epoxy composite skin. Its primary role would be to house the integrated pressure-fed oxygen/methane OMS and RCS system. The Block 2 baseline SM would have a total mass of 13,647 kg, of which over 9,200 kg would be propellant. In the baseline lunar scenario, this system would be used to perform rendezvous and docking with the Lunar Surface Access Module (LSAM) and Cargo LV upper stage in Earth orbit. The combined spacecraft would be put on a trans-lunar trajectory by the upper stage of the Cargo LV. The LSAM would brake the combined 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. The SM would be required to make any lunar orbit contingency plane changes needed prior to ascent of the LSAM (evidently about 300 m/s of delta-V was allocated for this purpose). Following docking of the LSAM with the CEV, the LSAM would be cast away, and the SM would perform trans-earth injection. On approach to earth, the SM would perform a final self-disposal maneuver following separation from the CM.

The SM would be equipped with an integrated pressure-fed propulsion system consisting of one 66.7-kN OMS main engine and twenty-four 445-N RCS thrusters. These engines would be common to the SM and the LSAM ascent stage. In the event of a late ascent abort off the CLV, the SM OMS engines could also be used for separating from the LV and either aborting to near-coastline water landings or aborting to orbit. The SM propellant tanks were sized to perform up to 1,724 m/s of OMS and 50 m/s of RCS Delta-V with the CEV CM attached and 15 m/s of RCS Delta-V after separation. The fuel tanks would be Al-Li graphite wrapped, rated at 22.1 atmospheres, and use gaseous helium pressurant.

Two deployable, single-axis gimbaling solar arrays would fold out from the SM to generate the necessary CEV power from Earth-Orbit Insertion (EOI) to CM-SM separation prior to entry. The solar arrays would use three-junction photovoltaic cells. Use of solar arrays instead of fuel cells eliminated the reactant mass requirements associated with providing power during the long dormant periods the ISS would have to spend attached to the ISS or waiting in lunar orbit. The solar arrays use state-of-the- art three-junction photovoltaic cells.

Four 7 m2 radiator panels were mounted on the exterior of the SM. These panels provided heat rejection capability for the CEV's ECLSS/ATCS (Environmental Closed Life Support System/Active Thermal Control System). The system used a 60% propylene glycol/40% water single phase fluid loop running from the CM, the interior of the SM, and the radiator panels.

Three versions of the SM would be used for resupply missions to the ISS. These versions of the SM would be identical to the Block 2 lunar exploration version, except that propellant would be off-loaded to reflect the lower Delta-V requirements of the ISS mission. The versions were:

The reference Mars mission utilized a Block 3 CEV to transfer a crew of six between Earth and an MTV at the beginning and end of the Mars exploration mission. The mission would require the SM to maneuver for two days in taking the CEV from a 55 km x 296 km initial orbit to a rendezvous and docking with the MTV in a circular orbit between 800 and 1200 km altitude. After the docking, the SM 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 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.

NASA decided to use the integrated pressure-fed LOX and methane OMS/RCS propulsion system for the SM based on performance and commonality with the ascent propulsion system on the LSAM. The risk associated with this type of propulsion for a lunar mission could be substantially reduced by developing the system early and flying it to the ISS. NASA recognized that there was a high risk in developing a LOX/methane propulsion system by 2011, but development schedules for this type of propulsion system had been studied and were in the range of hypergolic systems.

Orbital Storage: 180 days. RCS Coarse No x Thrust: 24 x 440 N. RCS specific impulse: 315 sec. RCS total impulse: 11,700 kgf-sec. Spacecraft delta v: 1,724 m/s (5,656 ft/sec). Electric System: 9.00 average kW.

Family: Manned spacecraft module, Moon. Country: USA. Spacecraft: Orion CEV. Propellants: Lox/LCH4. Agency: NASA.

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