Centrojet

Aerojet Nitric acid/Amine rocket engine. Work begun June 1943. Torque to rotate the main shaft of the propellant pumps was developed by the canted engine nozzles themselves at the aft end of the shaft. Abandoned 1945.

After the spectacular explosion and cancellation of the government contract for the Aerotojet, Chandler Ross was assigned to nevertheless continue to perfect the design in a smaller-scale version called the Centrojet. Work had begun on this in June 1943. The engine operated on the same principle as the Aerotojet - the torque to rotate the main shaft of the propellant pumps was developed by the canted engine nozzles themselves at the aft end of the shaft. The Centrojet shaft differed from the Aerotojet in that it terminated in a single thrust chamber with at first four, and finally two fixed nozzles projecting. This was expected to give the design dynamic balance, compared to the Aerotojet's swiveling separate combustion chambers.

The shaft drove a gear box which powered the propellant pumps at a more moderate speed than the Aerotojet. A high torque electric motor started the engine through the gear system. Unlike the Aerotojet, Centrojet development was successful and the resulting engine was a "delightful toy to work with". The engine was controlled simply by-passing fluid around the pumps to vary speed and torque. The principal problems remained an unexplainable and insoluble leakage of propellants at the pump and shaft transfer seals.

However tests showed that as the shaft speed increased, the overall specific impulse decreased faster than would be expected by losses due to the side components of thrust. The rotation was causing some kind of efficiency degradation in the combustion process, but this could never be understood. However the stable unit was run in a series of serious combustion studies. But tests showed that the rotating engine dynamic approach could not be scaled up to rocket engines of thousands of pounds of thrust. So the technology was abandoned in 1945.

Thrust (sl): 2.250 kN (506 lbf). Thrust (sl): 230 kgf. Propellant Formulation: RFNA/Aniline+Furfural Alcohol.