Encyclopedia Astronautica
Cobra



cobra.jpg
Cobra
Pratt and Whitney lox/lh2 rocket engine. 4500 kN. Design 2003. Proposed as a long-life, moderate-to high-thrust, reusable booster engine that incorporated a safe, low-cost, low-risk, LH2/LOX single burner, using a fuel-rich, staged combustion cycle.

In 2003 Pratt and Whitney and Aerojet teamed to provide a wide range of main propulsion options in support of NASA's conceptual second generation, reusable launch vehicles. Pratt and Whitney's experience in reusable rocket turbo-machinery and health and maintenance management systems used for jet engines complemented Aerojet's experience in combustion devices and the integration and production of large propulsion systems for the Titan family.

COBRA was proposed as a long-life, moderate-to high-thrust, reusable booster engine that incorporated a safe, low-cost, low-risk, LH2/LOX single burner, using a fuel-rich, staged combustion cycle. Mature, flight proven Space Shuttle Main Engine alternate turbopumps reduced technical, schedule, and programmatic risk while at the same time meeting safety and reliability goals.

Features of the engine included:

  • Single pre-burner, fuel-rich staged combustion cycle
  • Double containment, failsafe powerhead, hot gas system
  • Incorporated flight-certified SSME Block II turbopumps
  • Long-life, robust, milled-channel nozzle construction eliminated hot side weld joints
  • Smooth start transients avoided life-limiting thermal stress
  • Blanch-shielded, formed-platelet liner technology reduced hotwall stress
  • Turbine inlet temperature reduced 500°F relative to the SSME, promoted longer life
  • Integrated engine controls and health management system enhanced safety and maintainability
  • Low development cost/risk based on use of flight-qualified hardware and mature technologies
  • Thrust (vac): 200,000-1,000,000 lbf
  • Dry Weight (at 600k thrust): 8,000 lbm
  • Specific impulse (vac): 455 sec
  • Cycle: Staged combustion
  • Propellants: Liquid hydrogen / liquid oxygen
  • Mixture ratio: 5.5:1 to 6.5:1
  • Shutdown reliability: 0.9995
  • Catastrophic reliability: 0.999995
  • Mission life: > 100 missions
  • Time between overhauls: > 50 missions
  • Scheduled maintenance per flight: < 100 man-hours
  • Turnaround between flights: < 16 hours

An engine that could meet these specifications would be what what the SSME was originally designed to be. Whether this was technically possible was another matter…

Engine: 3,600 kg (7,900 lb). Oxidizer to Fuel Ratio: 6.

Status: Design 2003.
Unfuelled mass: 3,600 kg (7,900 lb).
Thrust: 4,500.00 kN (1,011,600 lbf).
First Launch: 2003.

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Associated Manufacturers and Agencies
Associated Propellants
  • Lox/LH2 Liquid oxygen was the earliest, cheapest, safest, and eventually the preferred oxidiser for large space launchers. Its main drawback is that it is moderately cryogenic, and therefore not suitable for military uses where storage of the fuelled missile and quick launch are required. Liquid hydrogen was identified by all the leading rocket visionaries as the theoretically ideal rocket fuel. It had big drawbacks, however - it was highly cryogenic, and it had a very low density, making for large tanks. The United States mastered hydrogen technology for the highly classified Lockheed CL-400 Suntan reconnaissance aircraft in the mid-1950's. The technology was transferred to the Centaur rocket stage program, and by the mid-1960's the United States was flying the Centaur and Saturn upper stages using the fuel. It was adopted for the core of the space shuttle, and Centaur stages still fly today. More...

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