Would it be possible to run the turbopump rich enough to inject enough gas into the engine bell to basically force the flow separation to happen at the turbopump exhaust injection site? If the exhaust vector was stable, and the engine bell wasn't shaking apart, it seems this change might make it possible to land an F9 US.
The exit pressure of an MVac is really low.
If you tried to land an F9 US, the engine nozzle would suffer flow separation.
Among other things, the thrust vector could shift around in an unpredictable manner.
The MVac injects the turbopump exhaust into the engine bell. Suppose, during the landing burn, you ran the MVac turbopump with a richer fuel:oxidizer ratio. For the same amount of pump power, you'd emit more gas.
Don't you think all those forces on the nozzle extension would rip it apart?
Quote from: cuddihy on 07/29/2016 08:57 pmDon't you think all those forces on the nozzle extension would rip it apart?No. As long as the stagnation pressure holds it under tension (at low densities before start). The concern is oxidation/heating nearing EI, thus the film cooling.The point in starting the engine as low as possible is to waste as little props as possible, because you need the drag as much/more than the delta-v, and that goes up with density.The weakness in this approach is the fact you need the engine's plume on the edges and not the center, so the angle of the shock wave flattens/broadens, thus the mention above of intentional turbulent flow generation above the top of the vehicle, acting as a drag chute. That's where things will get rough. What happens as the vehicle slows down, is that will approach the top of the vehicle - vehicle must go transonic before that occurs.Although the lack of the stiffener ring concerns after engine start for increasing densities out of concern that the TO of the engine and the retro plume chaotically leverage to destroy the nozzle extension. Probably for that reason you'll have to give up some iSP, but suggest one might be able to acoustically modulate combustion to bound the effect.[/quoteCouple things: The extension is radiatively cooled; radiating heat into the exhaust plume and surrounding opaque plasma during high velocity entry probably won't work well at all.For any sort of tail-first lifting entry (like the F9 S1) there will be asymmetrical aerodynamic loads on the nozzle extension, which WILL cause bending in the extension with resulting tension and compression loads. A tail-first ballistic entry might be able to keep the loads on the extension in tension only, but at the cost of much higher peak heating rates and drag loads.Based on these, I think retracting or discarding the extension are the only viable options for S2 retropropulsion in any atmosphere.
Couple things: The extension is radiatively cooled; radiating heat into the exhaust plume and surrounding opaque plasma during high velocity entry probably won't work well at all.
For any sort of tail-first lifting entry (like the F9 S1) there will be asymmetrical aerodynamic loads on the nozzle extension, which WILL cause bending in the extension with resulting tension and compression loads.
A tail-first ballistic entry might be able to keep the loads on the extension in tension only, but at the cost of much higher peak heating rates and drag loads.
Quote from: envy887 on 07/30/2016 03:24 pmCouple things: The extension is radiatively cooled; radiating heat into the exhaust plume and surrounding opaque plasma during high velocity entry probably won't work well at all.Consider flow rate and the fact that the entry heating is far hotter than exhaust. Also, with radiative heating, the IR passes through the less dense atmosphere and isn't backscattered - that is the reason why you can use such an extension in the first place on a US engine. Duh.
..You might be able to stabilize the stage post engine fire nozzle rearward and avoid damage. Passively/actively. Above dense atmosphere. If you care. That wasn't the point here. Like Mars. Mars doesn't have high density atmosphere.The point is using high altitude, high iSP to recover a US. And I can't "understand it for you" if you choose to use the wrong profile to evaluate the approach. Again, it is not a booster profile. Duh, its a US.
Jim, sigh, your answer doesn't seem to address my proposal. Running the turbopump richer will produce more turbopump exhaust for a given amount of thrust.
Terminal velocities vs altitude for a 4500 kg, 3.66 m diameter Cd=1 vehicle. The stage doesn't have Cd=1 but it will be close enough for this to be a guide.5 km 736 g/m^3 101 m/s10 km 413 g/m^3 135 m/s15 km 194 g/m^3 196 m/s20 km 88 g/m^3 293 m/s25 km 40 g/m^3 432 m/sHelicopter recovery is pretty close to sea level:CH-53E payload: 14,500 kgCH-53E service ceiling: 5640 m, 686 g/m^3Airplane recovery is just a bit better:C-17 payload: 77,520 kgC-17 service ceiling: 13716 m, 237 g/m^3An unmodified MVac will not be firing during capture conditions. But capture of the upper stage while under a parachute seems plausible. So that leaves a question of whether the engine can get the stage down to velocities compatible with a parachute.The erratic gusts of exhaust from the first stage during entry retroburn do not suggest static forces on the engine bell. This seems like a killer.Jim, sigh, your answer doesn't seem to address my proposal. Running the turbopump richer will produce more turbopump exhaust for a given amount of thrust. I was wondering if there is any guidance available on how much turbopump exhaust would be needed to make the MVac's thrust vector stable in a higher-pressure environment.
Well now THAT's interesting Jim. Thank you.They throttle the gas generator with fuel. The GG is run rich, so that means the heat output in the GG's burner is basically constant, and they throttle it by changing between low temp, high volume and high temp, low volume. I suppose it's a lot easier to throttle the non-cyrogenic fluid. And a given change to turbine output comes from a much larger change in fuel flow than oxygen flow, so that probably helps get finer control.
What about the GG exhaust injection enabling a clean, predictable flow separation from the main engine nozzle walls?
So that makes it seem like the diagram you have, which is for the Merlin 2, is just a different control system than for the Merlin 1.
which increases the pump pressure, which increases the fuel and oxidizer flow to both the main chamber and the GG, and that increases the GG gas volume which is positive feedback.....And there is one other thing. CRS-6 came down hard because of a "stiction in the biprop throttle valve, resulting in control system phase lag" (Elon Musk). This suggests they are throttling both LOX and fuel to something, I suppose the gas generator.
Okay, so both the RP-1 and LOX are modulated by throttle valves on the way into the gas generator. That means that during the landing burn it would be possible to run the gas generator richer than during the ascent burn, so as to increase the amount of turbopump exhaust volume injected into the engine bell for a given amount of turbopump power. Well, unless something is going to break from running at lower temperature (like maybe gas generator ignition?)
Two identical turbines. Both have the same mass flow entering them. One gas stream is hotter than the other. Output ambient pressure is equal.
The hotter turbine will produce more output power on the shaft, right?
Two identical turbines. Both have the same mass flow entering them. One gas stream is hotter than the other. Output ambient pressure is equal.The hotter turbine will produce more output power on the shaft, right?
As you say, PV=nRT. Turbine A's incoming stream has larger volume. Output stream will have larger volume. Shaft output power will be larger, because the shaft will be turning faster.
Two identical turbines. Both have the same mass flow entering them. Exit pressure is the same. Let's say the torque load on the shaft is the same. Turbine A has hotter gas. That's enough to determine the rest of the conditions.As you say, PV=nRT. Turbine A's incoming stream has larger volume. Output stream will have larger volume. Shaft output power will be larger, because the shaft will be turning faster.Right?
1. And that means I could reduce the mass flow for the hotter one to make the shaft power the same. And then I'd have two turbines with the same shaft power emitting different mass flows of gas. One emits less mass at higher temperature and lower volume. 2. This demonstrates that, by actuating the turbopump's LOX and RP-1 valves separately, we can vary the GG output mass rate for a fixed turbopump shaft power. We don't have full freedom, because the GG can't get too hot, but we can go some amount in the other direction and run the GG with more mass flow and lower temperature.
The F-1 engine pictures show that it's possible to get the GG exhaust well past the exit of the bell before it mixes with the main flow. .
The GG exhaust in the bell is subsonic, so it will decelerate and increase in pressure as it expands towards the exit of the bell.
So then once I can vary the mass flow from the GG into the main engine bell, my question was around how much control I can actually get from this.
.........And, supersonic gases accelerate when the flow cross section expands. Subsonic gases decelerate. That's why rocket nozzles coverge and then diverge. Now you could tell me that the GG exhaust is injected into the bell supersonically, and I'd be surprised, but I suppose it's possible. Short of that, though... I'm starting to think we're not actually having the conversation I was hoping to have.That's too bad. Oh well.
The shear between the subsonic GG exhaust (<500 m/s) and the main flow in the F-1 (>2600 m/s)
Mixing will be greater on the side with more shear.
Uh... are you just pulling my leg?
Now you could tell me that the GG exhaust is injected into the bell supersonically,
I'm starting to think we're not actually having the conversation I was hoping to have.
i just do not want to make another topic so this seems close enough.what if reuse just most valuable part of the 2 stage - engine? it has compact size so no need to big heat shield. it has integrated heat tolerant tail stabilizer so no need to worry about center of mass problems. it is light weighted (400 kg?) so no need for big shute an the end. and even small helicopter could catch it in the air.
Quote from: drzerg on 03/17/2017 04:37 ami just do not want to make another topic so this seems close enough.what if reuse just most valuable part of the 2 stage - engine? it has compact size so no need to big heat shield. it has integrated heat tolerant tail stabilizer so no need to worry about center of mass problems. it is light weighted (400 kg?) so no need for big shute an the end. and even small helicopter could catch it in the air. My gut feeling is that it won't pass cost-benefit analysis. You're up against the cost of relatively inexpensive engines Vs flight rate Vs the cost of trained recovery crews, recovery equipment, recovery operations, inspection, re-certification, and re-integration of the engines. It would probably make more sense to re-design Stage 2 around the ITS paradigm than perform airborne helicopter recovery of second stage hardware.
possible solution is that after deorbit burn engine could detach from stage and make its own additional deorbit "burn" with cold gas from main nozzle. so they could be separated at the point of impact.