> And ion thrusters have an exhaust velocity in the 20-50 km/s range
I compared with the Space Shuttle main engine because they are both hydrogen-oxygen engines.
As for the speed they mention, it's not clear why they focus on that one. At the end of the day, there are only 3 things that matter for a rocket engine: exhaust velocity, how fast it can burn the fuel, and the mass of the engine. You want the highest exhaust velocity, and the lowest mass of the engine. For the upper stages, you don't need to be able to burn fuel that fast, but you wouldn't mind that either.
All other things are implementation details. Now, the Space Shuttle main engine achieved 86% of the maximum theoretical exhaust velocity [1]. If this innovation here managed to increase this to 90% or more, that would be quite interesting, but hardly groundbreaking if it adds substantial mass to the engine.
> I'm interested to hear more about the possibilities this combustion chamber design brings up.
I would be interested too. This article was quite unilluminating unfortunately.
> At the end of the day, there are only 3 things that matter for a rocket engine: exhaust velocity, how fast it can burn the fuel, and the mass of the engine.
While those are primary factors, and a rocket won't go up if they aren't good enough, they aren't the only ones (or else every rocket engine would be hydrolox staged combustion, like the SSMEs). Other key factors are engine cost (SSME does terrible here), fuel type (hydrolox needs way larger diameter boosters, isn't worth it for first stage), combustion stability (allows engines to throttle, key for max-Q and reducing max g as stage empties, especially on upper stage), reusability (important for some systems), controllability (how quickly it can throttle)... While their importance gradually diminishes, there is a list hundreds long of different trade-offs made in engine development. The reason I said I'd like to know more is indeed because that article wasn't informative. I certainly am not writing it off because it can't be plopped in a rocket right now
Thermodynamic efficiency matters. And detonation is highly efficient, because it's an (almost) combustion isochoric process. The shockwave/detonation front travels at far higher speeds than the movement/speed-of-sound in both the unburnt input mix, as well as the combustion products.
My understanding is that it's due to the combustion performing all it's work before any work/energy is extracted from the hot combustion products. This results in a higher peak temperature, and the higher the temperature difference, the higher the theoretical efficiency limit of converting the thermal energy into mechanical (or equivalent) energy. It's why diesel engines have a lower fuel consumption than gasoline engines: they burn hotter.
I compared with the Space Shuttle main engine because they are both hydrogen-oxygen engines.
As for the speed they mention, it's not clear why they focus on that one. At the end of the day, there are only 3 things that matter for a rocket engine: exhaust velocity, how fast it can burn the fuel, and the mass of the engine. You want the highest exhaust velocity, and the lowest mass of the engine. For the upper stages, you don't need to be able to burn fuel that fast, but you wouldn't mind that either.
All other things are implementation details. Now, the Space Shuttle main engine achieved 86% of the maximum theoretical exhaust velocity [1]. If this innovation here managed to increase this to 90% or more, that would be quite interesting, but hardly groundbreaking if it adds substantial mass to the engine.
> I'm interested to hear more about the possibilities this combustion chamber design brings up.
I would be interested too. This article was quite unilluminating unfortunately.
[1] https://space.stackexchange.com/questions/17129/is-there-a-m...