Status: Canceled 1960. Payload: 1,400 kg (3,000 lb). Thrust: 667.20 kN (149,993 lbf). Gross mass: 53,000 kg (116,000 lb). Height: 30.00 m (98.00 ft). Diameter: 2.44 m (8.00 ft). Apogee: 185 km (114 mi).
The Ablestar complete upper stage under the leadership of Tiger Eldridge at Space General was a resounding success, and when Aerojet became the Mission Planning Contractor for the Thor Ablestar satellite launch vehicle, Aerojet began to look forward to the next step, a high energy version of Ablestar. Studies indicated that a Liquid Hydrogen/Liquid Oxygen design was the logical choice.
An outside interest lent impetus to this effort when John Kuhn of STL suggested that the use of LOX/LH2 and ablative materials like those being considered for re-entry vehicles might be appropriate for thrust chambers. Also, STL/Air Force had earlier funded some development work on an ablative thrust chamber as part of the Ablestar production contract. This was carried through completion of qualification testing, but was never used in production. At about the same time, the ablative chamber for SAINT was being considered. An Aerojet sponsored effort to evaluate the concept was undertaken in Azusa, combining all this with the general liquid propellant background, cryogenic engineering experience available from the Rover program, earlier LOX/LH2 development work, and the knowledge and fabrication capability resident in the Structural Plastics organization.
A thrust chamber assembly was made using ablative material in tape form laid by hand on a mandrel compatible with a Titan II second stage injector adapted for LOX/LH2. This included redrilling the injector face and inserting a gas generator igniter from the Bell Aircraft. Rascal engine as a starter. An initial test resulted in a burn-out in little over 5 seconds, which was found to have been caused by an improper choice of method of fabrication of the chamber section. With the proper design and fabrication, a later unit was test fired at the Sacramento test facility using pressure fed propellants for a duration of 193 seconds. The ablative material showed remarkably little damage, and the test duration was limited only by the amount o: hydrogen available.
Test results and specimens cut from the chamber created considerable interest in the Air Force. A subsequent briefing on the Aerojet concept of a LOX/LH2 Upper Stage using pressure-fed propellants and an ablative thrust chamber led to Contract AF 04 (611) 5170. This effort included experimental and design studies. Flight-type thrust chambers were fabricated and proven at operating conditions. The flight pressurization concept was demonstrated at quarter scale. Design studies established the size and configuration of the stage, and mission analysis showed that such an upper stage could perform a wide range of orbital missions using existing ballistic missiles as boosters. The design was called the Hydra after the many headed creature from Greek mythology.
The Air Force elected not to continue the Hydra concept, partly because of the sensitivity to the relatively recent decision to have NASA handle all new space launch vehicles - and also quite likely because of the availability of Agena, and the higher performance Centaur project using more complex pump-fed Pratt & Whitney RL-10 LOX/LH2 engines which were in a relatively more advanced stage of development.
This hydrogen work on Hydra was conducted in approximately the same 1959-1960 time frame as the little known Air Force LOX/LH2 program involving conversion of the Titan I first stage to use these propellants by Aerojet Sacramento.
LEO Payload: 1,400 kg (3,000 lb) to a 185 km orbit at 28.00 degrees.