The LOX/CH4 engine described in the paper EXPERIMENTAL STUDIES
ON COMBUSTION INSTABILITIES
IN A MULTI-ELEMENT, RECTANGULAR ROCKET CHAMBER is stated to have an L* between 113–339 cm.
The paper Comprehensive Design Method for LOX/Liquid-Methane Regenerative
Cooling Combustor with Coaxial Injector gives a Ts between 0.97 and 1.71 ms.
Radial burn is usually preferred.
Axial burn rockets (solid rockets that burn from the bottom up – otherwise known as end burners) suffer from the problem that the centre of gravity of the rocket changes as the rocket burns.
Radial burn rockets do not tend to suffer from this problem as they mostly burn from the inside out along the length of the rocket.
All commenters, don't mix chemical oxygen reactions with oxygen production and depletion. Photosynthesis does not produce oxygen atoms, also breathing and burning of fossil fuels does not consume. Its just a transition because oxygen reactions converts chemical energy in the following example equations [CH4 + O2 <-> CO2 + H2O + Energy] as Earth green ...
I noticed between the info in the question, additional info in comments, and an assumption of 0.0 delta thrust at 0.0 flowrate, there was enough information to draw a exponential curve for the throttle.
Delta Thrust (N)
Inert Prop Flowrate (kg/s)
Using this throttle characteristic curve, I ...
Reaction Control System Thrusters typically have a thrust on the order of Newtons to tens of Newtons – hundreds of Newtons for atmospheric control (e.g. the RCS thrusters on Falcon 9 have about 400 N). Ion Thrusters in current use have a thrust on the order tens to hundreds of Millinewtons. There are experimental thrusters that have achieved 5 N in lab ...
The RL10 has been experimentally fired on methane and propane as well as hydrogen. This did entail modifications to several components of the engine; for the methane version:
Fuel component modifications were unique and were accomplished under the contract; assemblies that were changed were the pump, turbine, thrust control, injector, and thrust chamber.
Let's look at A11 for simplicity. Some key points from Apollo by the Numbers, p 307, rounded a bit:
"Empty" LM at undocking - 5460 lbs (including everything left on board)
CSM at start of TEI - 36970 lbs (including everything on board)
CSM at cut-off - 26790 lbs (10180 lbs used)
The mission report, p 99, gives the overall propellant values, again ...
In the realm of theoretically possible but thoroughly impractical, consider that magnetic fields are ubiquitous in space. Charged particles as engine exhaust will follow helical trajectories. Put your collector at a point where the exhaust has traversed half a helical cycle. This transfers momentum to the magnetic field, yielding thrust without expending ...
Not practical. There are several problems.
First, what comes out of a rocket's rear end is mostly not even fuel.
Second, the big cloud of fuel you see upon launch is mostly water that prevents the rocket from tearing itself apart with its own shockwaves.
Last, but the most importantly, even if the exhaust was actually a new fuel that combusted with its parts ...
If you ever come up with a chemical solution where re-burning exhaust looks look like a good idea, don't build this. It won't work because catching the exhaust cancels out the fuel consumption. It's easier to see when thinking of momentum rather than energy. You've got to leave reaction mass behind.
On the other hand, if you find the chemical solution, the ...