Encyclopedia Astronautica
Vulcain


SEP, Ottobrunn lox/lh2 rocket engine. 1075 kN. In production. Isp=431s. Powered the cryogenic core stage of Ariane 5. First flight 1996.

The Vulcain rocket engine powered the cryogenic core stage of Ariane 5. The Ottobrunn facility was responsible for the development and manufacture of the Vulcain thrust chamber comprising the regeneratively cooled combustion chamber, coaxial propellant mixing injectors, dump cooled nozzle extension, and gimbal joint. The LOX and LH2 propellant valves were also manufactured and produced at the Ottobrunn Production Centre.

The thrust chamber design was based on the regenerative cooling concept that was developed at Ottobrunn, since continually refined. Before combustion, LH2 was pumped into a distribution manifold and then flowed through closely arranged small tubular cooling channels within the combustion chamber wall. The LH2 then entered an injector head where it was uniformly distributed to 516 coaxial injector elements.

The coaxial injector elements caused the LOX and LH2 propellants to be mixed together. LOX was injected at the centre of the injector, around which the LH2 was injected. These propellants were mainly atomized and mixed by shear forces generated by the velocity differences between LOX and LH2. Althought the injector design was complex, it assured consistent and reliable combustion eficiencies greater than 99 %, which were reached in the remaining process in the combustion chamber.

At the combustion chamber, the mixed propellants were burned and accelerated up to sonic conditions. The combustion temperatures in the chamber almost reached 3250 degrees Celcius at pressures greater than 100 bar.

Combustion temperature control was achieved by the flow of LH2 in the cooling channels within the combustion chamber wall. This thin copper alloy wall, just 1.5 mm thick separated the combustion temperatures from the - 239 to - 120 degree Celcius LH2 cooling flow.

The final acceleration of hot gases, up to supersonic velocities, was achieved by gas expansion in the nozzle extension, thereby increasing the thrust.

Thrust (sl): 773.200 kN (173,822 lbf). Thrust (sl): 78,844 kgf. Engine: 625 kg (1,377 lb). Chamber Pressure: 102.00 bar. Area Ratio: 45. Thrust to Weight Ratio: 84.318. Oxidizer to Fuel Ratio: 6.2. Coefficient of Thrust vacuum: 1.86973284462013. Coefficient of Thrust sea level: 1.4411614160487.

Status: In production.
Unfuelled mass: 1,300 kg (2,800 lb).
Height: 3.00 m (9.80 ft).
Diameter: 2.00 m (6.50 ft).
Thrust: 1,075.00 kN (241,669 lbf).
Specific impulse: 431 s.
Specific impulse sea level: 326 s.
Burn time: 605 s.
Number: 22 .

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Associated Countries
See also
Associated Launch Vehicles
  • Ariane 5 French orbital launch vehicle. The Ariane 5 was a completely new design, unrelated to the earlier Ariane 1 to 4. It consisted of a single-engine Lox/LH2 core stage flanked by two solid rocket boosters. Preparatory work began in 1984. Full scale development began in 1988 and cost $ 8 billion. The design was sized for the Hermes manned spaceplane, later cancelled. This resulted in the booster being a bit too large for the main commercial payload, geosynchronous communications satellites. As a result, development of an uprated version capable of launching two such satellites at a time was funded in 2000. More...

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...

Bibliography
  • Kudryavtseva, V M, ed., Zhidkostnikh Raketnikh Dvigatley, Visshaya Shkola, Moscow, 1993.

Associated Stages
  • Ariane H155 Lox/LH2 propellant rocket stage. Loaded/empty mass 170,800/12,700 kg. Thrust 1,114.00 kN. Vacuum specific impulse 430 seconds. Chamber pressure 108 bar; expansion ratio 45.0; propellant mix ratio 5.3. More...

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