Space Launch System (SLS)

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The Space Launch System (SLS) is an American Space Shuttle-derived heavy expendable launch vehicle being designed by NASA. It follows the cancellation of the Constellation program, and is to replace the retired Space Shuttle. The NASA Authorization Act of 2010 envisions the transformation of the Constellation program's Ares I and Ares V vehicle designs into a single launch vehicle usable for both crew and cargo, similar to the Ares IV. SLS will be the world's most powerful rocket with 20% more thrust than Saturn V and a comparable payload capacity.

The SLS launch vehicle is to be upgraded over time with more powerful versions. Its initial Block 1 version is to lift a payload of 70 metric tons to low Earth orbit (LEO), which will be increased with the debut of Block 1B and the Exploration Upper Stage. Block 2 will replace the initial Shuttle-derived boosters with advanced boosters and is planned to have a LEO capability of more than 130 metric tons to meet the congressional requirement. These upgrades will allow the SLS to lift astronauts and hardware to various beyond-LEO destinations: on a circumlunar trajectory as part of Exploration Mission 1 with Block 1, to a near-Earth asteroid in Exploration Mission 2 with Block 1B, and to Mars with Block 2. The SLS will launch the Orion Crew and Service Module and may support trips to the International Space Station if necessary. SLS will use the ground operations and launch facilities at NASA's Kennedy Space Center, Florida.

During the joint Senate-NASA presentation in September 2011, it was stated that the SLS program has a projected development cost of $18 billion through 2017, with $10 billion for the SLS rocket, $6 billion for the Orion Multi-Purpose Crew Vehicle and $2 billion for upgrades to the launch pad and other facilities at Kennedy Space Center.

Note that SLS specifications vary by date and source. Numbers are likely to change as design and development progresses.


Principal uses: Orion MPCV and cargo to deep space
LEO (48 × 296 km, 28.5o): 70,000 kg (Block 1), 105,000 kg (Block 1B), 130,000 (Block 2)
Translunar injection: 24,500 kg (Block 1), ≈45,000 kg (Block 1B), ≈60,000 (Block 2)
Cost: US$7 billion 2014-2018 (2014 estimate) to $35 billion until 2025 (2011 estimate)
Number of stages: 2 + 2 strap-ons
Overall length: 98 to 111 m (including payload)
Principal diameter: 8.4 m
Launch mass: 2,600 to 2,950 t
Launch thrust: 39 MN (Block 1, 1B), 41 MN (Block 2)
Contractors: Boeing, United Launch Alliance, Orbital ATK, Aerojet Rocketdyne

Engine: 5 segment SRM
Length: 53.9 m, including nose cone
Diameter: 3.71 m
Total mass: each 731,885 kg
Propellant: PBAN solid
Propellant mass: each 631,495 kg
Thrust: each 16,000 kN maximum
Specific impulse: 268 s vac
Burn time: 126 s

Engine: 4 segment SRM
Total mass: each ≈793,000 kg
Propellant mass: each ≈709,000 kg
Thrust: each 20,000 kN maximum
Specific impulse: ≈286 s vac
Burn time: 110 s

Engine: 4 RS-25D/E
Length: 64.6 m
Diameter: 8.4 m
Dry mass: 85,275 kg
Burnout mass: ≈112,000 kg
Oxidizer: liquid oxygen
Fuel: liquid hydrogen
Propellant mass: 979,452 kg (useable)
Thrust: 7,440 kN SL, 9,116 kN vac
Burn time: 476 s

Engine: 1 RL10B-2
Length: 13.7 m
Diameter: 5.0 m
Dry mass: 3,765 kg + ≈5,000 kg aft interstage
Burnout mass: 4,354 kg
Oxidizer: liquid oxygen
Fuel: liquid hydrogen
Propellant mass: 26,853 kg (useable)
Thrust: 110.1 kN vac
Burn time: 1125 s

Engine: 4 RL10-C
Length: not to exceed 18 m
Diameter: 8.4 m
Oxidizer: liquid oxygen
Fuel: liquid hydrogen
Propellant mass: up to 129,000 kg
Thrust: 440 kN vac

Diameter: 5.0 m (Block 1, 1B), 8.4 m (Block 1B), 10.0 m (Block 2)
Length: 19.1 m (Block 1, 1B), 31.1 m (Block 2)
Mass: about 4 to 12 tonne

Application: SLS core stage, until supply runs out
Dry mass: 3,525 kg
Engine cycle: staged combustion
Oxidizer: liquid oxygen, delivered at 439 kg/s
Fuel: liquid hydrogen, delivered at 73 kg/s
Mixture ratio: 5.85 to 6.10 ±1%
Thrust: 2,274 kN vac at max. power, can be throttled to 65-109%
Specific impulse: 452.3 s vac

Application: SLS core stage
Dry mass: not to exceed 3,700 kg
Engine cycle: staged combustion
Oxidizer: liquid oxygen, delivered at 445 kg/s
Fuel: liquid hydrogen, delivered at 74 kg/s
Mixture ratio: 5.85 to 6.10 ±1%
Thrust: 2,294 kN vac at max. power, can be throttled to 65-111%
Specific impulse: 450.8 s vac

Application: Interim cryogenic propulsion stage
Dry mass: 301 kg
Engine cycle: expander
Oxidizer: liquid oxygen, delivered at 20.6 kg/s
Fuel: liquid hydrogen, delivered at 3.5 kg/s
Mixture ratio: 5.88
Thrust: 110.1 kN vac
Specific impulse: 462 s vac

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