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Orion is the next generation crew piloted spacecraft for the United States. Planning began with the Vision for Space Exploration announcement in January 2004, with full-scale development commencing with the selection of Lockheed Martin as the prime contractor on 31-August-06. Four versions of the spacecraft are planned to complete the following missions: crew transport to the ISS in low earth orbit, unmanned pressurized cargo to the ISS, manned lunar flights, and manned Mars flights. Orion targets the first mission to the ISS no later than 2014 and to the Moon no later than 2020.
International Space Station (ISS) missions will be able to transport up to six crew members for ISS crew rotation. Orion will be able to stay in orbit for up to 210 days and serve as an emergency lifeboat for the entire ISS crew. For lunar missions, Orion will support a crew of four astronauts while in transit to and from the Moon and will serve as the Earth reentry vehicle. While lunar surface operations are ongoing, Orion will remain uninhabited in lunar orbit. For Mars missions, Orion will transport the crew to low Earth orbit and serve as the reentry vehicle. All Orion versions will be launched on the Ares I launch vehicle.
Orion consists of four major system elements: the launch abort system, the crew module, the service module, and the spacecraft adapter. The launch abort system is used for emergency escape during launch, the crew module is used for crew and cargo transport, the service module provides propulsion, electrical power and fluids storage, and the spacecraft adapter is the structural transition to the launch vehicle. The only system to return to Earth is the crew module, which will be partly reusable. Orion will be capable of either land or water landings.
Below are the preliminary specifications for Orion as released by NASA on 31-August-06.
ORION SPECIFICATIONS
First manned flight: planned for Sep-2014
Principal uses: manned earth and lunar orbiter
Cost: $3.9 billion for DDT&E and first unit, up to $3.5 billion for additional units
Crew size: up to 6
Orbital storage: 210 days
Principal diameter: 5.03 m
Habitable volume: 10.22 m3
Primary engine thrust: 33.36 kN
Main engine propellant: NTO/MMH
Total spacecraft delta v: 1,855 m/s
Power: solar cells, 9.15 kW
Contractor: Lockheed Martin
LAUNCH ABORT SYSTEM
The launch abort system consists of a nose cone, attitude control motor, canard section, jettison motor, interstage, abort motor, adapter cone, and boost protective cover.
Abort motor
nozzles: 4 each, canted 30o
specific impulse: 250 s
thrust: 2,252.6 kN (total in vehicle axis)
burn time: 2.0 s
thrust/weight ratio: 15:1
Attitude control motor:
nozzles: 8 each, canted 90o
specific impulse: 227 s
thrust: 11.12 kN
burn time: 20 s
Jettison motor:
nozzles: 4 each, canted 35o
specific impulse: 221 s
thrust: 43.01 kN
burn time: 1.5 s
Dry mass: 3,696 kg
Propellant mass: 2,480 kg
Launch mass: 6,176 kg
CREW MODULE
Crew size: up to 6 to ISS, 4 to Moon
Length: 3.30 m
Maximum diameter: 5.03 m
L/D hypersonic: 0.34 @ 157o
Pressurized volume: 19.59 m3
Habitable volume: 10.22 m3
Reaction control system
propellant: GOX/methane
delta v: 50 m/s
thrusters: 12 x 445 N
Dry mass: 7,891 kg
Propellant mass: 174.7 kg
Oxygen/nitrogen/water mass: 128.3 kg
Landing mass: 7,337 kg
Launch mass: 8,485 kg
Lunar return payload: 100 kg
SERVICE MODULE
Principal diameter: 5.03 m
Structural configuration: 3 rings / 6 longerons
Propulsion configuration: 2 X 2 serial feed
Propulsion system
propellant: NTO/MMH
delta v: 1,855 m/s
main engine thrust: 33.36 kN
RCS thrusters: 24 x 445 N
Solar array area: 36.0 m3
Solar array power: 9.15 kW
Radiator area: 31.0 m3
Thermal dissipation: 6.3 kW
SPACECRAFT ADAPTER
The spacecraft adapter is constructed of aluminum longerons and ring frame, with IM7/977-3 sandwich panel material.
Diameter: truncated cone tapering from 5.03 m to 5.50 m
Length: 3.31 m
Mass: 581 kg