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9

There is another important variable besides chamber pressure and nozzle design to consider--how the engine pumps the propellants into the chamber. The J-2 engine, like the lower stage (and much more powerful but lower ISP kerosene burning F-1 engine) used a gas generator (basically burning oxygen and fuel, here hydrogen) to drive a turbine that pumped the ...


9

Designed for high altitude low ambient pressure, I assume. The primary optimization is in the nozzle: For optimal performance the pressure of the gas at the end of the nozzle should just equal the ambient pressure: if the exhaust's pressure is lower than the ambient pressure, then the vehicle will be slowed by the difference in pressure between the top of ...


5

Remember when the J-2X engine was proposed for the Ares I and V rockets? That featured a nozzle area ratio of 55:1. This ratio was the same as that of the original J-2X proposal. However, unlike the new design that would add performance through gas generator and chamber enhancements, the old J-2X was a simplification study. The modern J-2X and its 55:1 ...


5

SSME wouldn't have been insane. RL-10 is used on relatively lightweight upper stages: Centaur and DCSS plus their payloads are in the 20-40 ton range, and the thrust-to-weight ratio winds up being in the ballpark of 0.3:1 -- not enough to lift off, but reasonable for circularizing your orbit once the first stage gets you most of the way to altitude. If you'...


5

A number of factors contributed to the complexity. Staged combustion was definitely a big one. The SSME also ran at a much higher chamber pressure than the J-2, 3000 psi vs 760 psi. This seems to have required two turbopumps for each of the propellants instead of the usual one -- there are low- and high-pressure turbopumps for the fuel (LPFTP, HPFTP) and low-...


5

The J-2 was not suited to firing at sea level. A sea level specific impulse of 200 seconds is widely quoted for the engine, but I'm suspicious of such a round figure; the exit pressure of the engine might be too low to fire at sea level without flow separation. Assuming that it could be fired at sea level, or air-started, a scaled-down version of the ...


2

It's really an apples to oranges comparison. J-2s are upper stage, gas generator, single-use engines. SSMEs are booster, staged combustion, reusable engines. Trying to compare them in that way is really fruitless. They were not designed to perform the same function.


2

FWIW, here are J2-S numbers. The paper "Altitude Developmental Testing of the J2-S" gives the throat diameter as 12.192 inches and the expansion ratio as 39.62. Paragraph 2.1.1 Thrust Chamber — The tubular-walled, bell-shaped thrust chamber consists of an 18.6-in.-diam combustion chamber with a throat diameter of 12.192 in., a characteristic length ...


2

This (which I found here) says 90 inches throat to exit. Two other items below show gimbal center to exit of 116 inches. So far I can not find the relationship between the throat and the gimbal center. above: Throat to exit - 90 inches. found here. above: Gimbal center to exit - 116 inches. See the "J-2S Side Drawing" link here. above: Gimbal center to ...


2

Rumor had it that the air-start of the SSME at stage separation was doable, but restarting it for any kind of orbital burn was not. I have never seen this openly discussed, and can hardly believe it wasn't addressed in the conceptual design, but something made them switch to J-2X with the subsequent dire effects on the program.


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