H-1 (rocket engine)
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The H-1 engine is a 200,000 lbf (890 kN) thrust LOX/RP-1 engine, used alone in the first stages of some Thor - Delta rockets and the Jupiter rocket. It is derived from the Navaho missile, and was simplified and improved for use in the S-IB first stage of the Saturn I and Saturn IB rockets. It is used in clusters of eight on all S-IB rocket stages. Later it would be uprated to 205,000 lbf (912 kN) of thrust. The H-1 preceded the F-1 engine, which was used on the Saturn V rocket.
Unlike the J-2 engine used on the S-IVB stage, the H-1 was a single-start engine. It could be fired multiple times -- and engines were usually subject to two or more static test firings before a mission to flight-qualify them -- but it could not be restarted in flight, because some components required for the startup sequence were non-reusable. In particular, the engine was ignited by a Solid Propellant Gas Generator (SPGG), which was essentially a small solid rocket, and had to be replaced after each firing.
To start the engine a 500V AC voltage was applied to the SPGG, which ignited the solid propellant. This produced hot gas which was allowed to build up until reaching a pressure of 600-700 psi, after which a bursting diaphragm released it into the turbine which drove the fuel turbopumps. This began the process of pumping fuel and oxidiser into the engine, and the hot gases from the SPGG provided the initial energy required to ignite the fuel/oxidizer mix. Once the fuel and oxidizer were being pumped and burning, the process was self-sustaining until engine shutdown.
[edit] Specifications
| Vehicle effectivity | ||
|---|---|---|
| SA-201 through SA-205 | SA-206 and subsequent | |
| Thrust (sea level) | 200,000 lbf (890 kN) | 205,000 lbf (912 kN) |
| Thrust duration | 155 s | 155 s |
| Specific impulse | 289 s | 289 s |
| Engine weight dry (inboard) | 1,830 lb (830 kg) | 2,200 lb (998 kg) |
| Engine weight dry (outboard) | 2,100 lb (953 kg) | 2,100 lb (953 kg) |
| Engine weight burnout | 2,200 lb (998 kg) | 2,200 lb (998 kg) |
| Exit-to-throat area ratio | 8:1 | 8:1 |
| Propellants | LOX & RP-1 | LOX & RP-1 |
| Mixture ratio | 2.23±2% | 2.23±2% |
| Fuel flow rate | 2092 USgal/min (132 L/s) | |
| Oxidizer flow rate | 3330 USgal/min (210 L/s) | |
| Nominal chamber pressure | 633 psia (4.4 MPa) | |
- Contractor: NAA/Rocketdyne
- Vehicle Application: Saturn I / S-IB 1st stage - 8-engines
- Vehicle Application: Saturn IB / S-IB 1st stage - 8-engines
[edit] References
Skylab Saturn IB Flight Manual, 30th September 1972
