Author Topic: Landing an MVac  (Read 9236 times)

Offline Jim

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Re: Landing an MVac
« Reply #20 on: 08/01/2016 06:40 PM »
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.


No, they are the same engine as far as that.  The differences are in the skirt extension and restart hardware.

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.


No, the CG valves (notice the "S" and the "bi", fuel is the control but ox is adjusted to keep about same mixture ratio) would modulate that and prevent the feedback.
« Last Edit: 08/01/2016 06:58 PM by Jim »

Offline IainMcClatchie

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Re: Landing an MVac
« Reply #21 on: 08/01/2016 11:03 PM »
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?)

But I haven't found any support for the notion that injected turbopump exhaust might encourage main combustion chamber gas to separate cleanly from the side of an overexpanded nozzle.

Offline Jim

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Re: Landing an MVac
« Reply #22 on: 08/02/2016 12:53 AM »
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?)

No, changing the mass flow in the gas generator is going change turbo pump output.  There is no way of changing GG output without affecting the turbo pump.

And the mass flow in the nozzle has little effect on flow attachment. 


Offline IainMcClatchie

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Re: Landing an MVac
« Reply #23 on: 08/02/2016 09:53 PM »
Uh oh, maybe there is something super basic that I'm missing.

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?

Offline CameronD

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Re: Landing an MVac
« Reply #24 on: 08/03/2016 04:03 AM »
Two identical turbines.  Both have the same mass flow entering them.  One gas stream is hotter than the other.  Output ambient pressure is equal.

Since no-one else has chipped in, I'll have a go.. :)

IIRC (and someone please correct me if I'm wrong), mass flow of a gas is related to volume flow rate, density and temperature so the only way to get a condition where the mass flow is the same but one stream is hotter is to reduce the density and/or the volume flow rate of the gas stream to match.  Eg. In a combustion situation, hotter generally means more fuel input and higher density, but to move away from stoich means less power.. so to answer your question:

The hotter turbine will produce more output power on the shaft, right?

AIUI, no, because for the same mass flow rate input the gas output would have to be slower.
   
« Last Edit: 08/03/2016 11:48 PM by CameronD »
With sufficient thrust, pigs fly just fine - however, this is not necessarily a good idea. It is hard to be sure where they are
going to land, and it could be dangerous sitting under them as they fly overhead.

Offline IainMcClatchie

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Re: Landing an MVac
« Reply #25 on: 08/03/2016 04:12 AM »
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?

Offline Dante80

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Re: Landing an MVac
« Reply #26 on: 08/03/2016 05:58 AM »
A strictly hypothetical question. It has been said above that the nozzle extension for MVac is very fragile and would shatter at re-entry. Also, the very large expansion ratio doesn't help with downthrottling stability.

What would be the hypothetical result if the radiative nozzle extension was dropped before re-entry happened?

Offline IntoTheVoid

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Re: Landing an MVac
« Reply #27 on: 08/03/2016 06:07 AM »
(not a rocket scientist)
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?

Why would you presume they have the same mass flow, rather than volume?
Why would you presume the output pressure is equal?
How would the pump be 'generating' power, rather than the gas generator?
What I see is the gas generator driving a common shaft to a fuel pump and oxidizer pump, which may not be sized identically. The fuel line then goes through a trim valve and then a MFV which I presume to be a Main Fuel Valve. The oxidizer line then goes to a MOV, which I presume to be a Main Oxidizer Valve.
The pressure in the lines would be based on all of these things, but not necessarily equal in mass, volume or pressure. The only thing I see required to be equal between fuel and oxidizer is the pump shaft rotational rate.

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.

How would the shaft be turning faster than itself?
I could be way off base, but I don't see where you're getting any of your assumptions from the diagram.

Offline Jim

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Re: Landing an MVac
« Reply #28 on: 08/03/2016 03:57 PM »
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?


Can't say exit pressure is the same.  Work was done. 

Still don't see what you are getting at?  There isn't a way to separate GG output from thrust level.  If you want to prevent flow separation from a large nozzle at lower thrust level, a secondary source, separate from the GG is going to have to supply the fluid.  And that still is not likely going to solve the issue, since the secondary fluid injection is going to generate thrust, further exasperating the issue.
« Last Edit: 08/03/2016 04:04 PM by Jim »

Offline IainMcClatchie

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Re: Landing an MVac
« Reply #29 on: 08/03/2016 08:08 PM »
I'm trying to strip out as many things as possible so we can get to something that we agree on.  IntoTheVoid, that means I'm comparing two different operating points for the turbine that powers the pumps.  I'm not comparing the two pumps.

If both turbines are exhausting to the same ambient, then exit pressures are the same.  Right?

Now I have two turbines producing different shaft power with the same mass flow, because one is running hotter than the other.  Right?

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.  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.

That's not even my main point, that's just a lemma that I figured everyone would agree on because it was obvious.  I've been surprised by how hard it is to communicate.  Sorry about not being more clear earlier.

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.  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.  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.  It seems like this might be a way to make a variable expansion nozzle for a GG type engine.

Offline Jim

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Re: Landing an MVac
« Reply #30 on: 08/03/2016 08:19 PM »

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.


1.  Not necessarily.  It is not automatic.  Turbine design has a play in it.

2.  And no, there are other issues with using more fuel such as coking and sustain combustion. 
« Last Edit: 08/03/2016 08:30 PM by Jim »

Offline Jim

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Re: Landing an MVac
« Reply #31 on: 08/03/2016 08:20 PM »
  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. .

No, it isn't mixing with the main flow, it is mixing with ambient air.  That is why it ignites.

  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.

It is decreasing in pressure, that is what the bell does.


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.

Little to nothing.  It is a small fraction of the engine output and adjusting it as you propose will have little effect on the overall engine thrust
« Last Edit: 08/03/2016 08:28 PM by Jim »

Offline IainMcClatchie

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Re: Landing an MVac
« Reply #32 on: 08/04/2016 12:03 AM »
Uh... are you just pulling my leg?

The shear between the subsonic GG exhaust (<500 m/s) and the main flow in the F-1 (>2600 m/s) is vastly greater than the shear between that GG exhaust and the ambient air.  Mixing will be greater on the side with more shear.  This ain't subtle.

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.

Offline CameronD

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Re: Landing an MVac
« Reply #33 on: 08/04/2016 02:10 AM »
.........
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.

Well, it's an interesting conversation nonetheless.. so keep it up. :)

If Jim is saying what I think he is, ISTM, in summary, that (a) you can't achieve what you're hoping to achieve without a complete re-design of the engine plumbing and (b) the GG exhaust isn't going to cut it (pardon the pun).

How about other proposals - like installing a cluster of smaller nozzles (say 4?) around the outside of the main nozzle and using those for landing instead?
« Last Edit: 08/04/2016 02:11 AM by CameronD »
With sufficient thrust, pigs fly just fine - however, this is not necessarily a good idea. It is hard to be sure where they are
going to land, and it could be dangerous sitting under them as they fly overhead.

Offline Jim

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Re: Landing an MVac
« Reply #34 on: 08/04/2016 02:38 AM »

The shear between the subsonic GG exhaust (<500 m/s) and the main flow in the F-1 (>2600 m/s)

Where do you get these numbers?
Mixing will be greater on the side with more shear.

Based on what principle?

Uh... are you just pulling my leg?

No, pressure decreases from the throat to the nozzle exit.

  Now you could tell me that the GG exhaust is injected into the bell supersonically,

It is injected at a point that it is at a higher pressure than in the nozzle at that point.   And the flow stays along the nozzle extension to cool it.  Look at the Merlin diagram.  Turbopump exhaust is less than 100 psi and the thrust chamber is at over 1400 psi.
« Last Edit: 08/04/2016 02:46 AM by Jim »

Offline Jim

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Re: Landing an MVac
« Reply #35 on: 08/04/2016 02:49 AM »
I'm starting to think we're not actually having the conversation I was hoping to have.


Which is the GG output can't be controlled independently of thrust.  Nor can turbopump exhaust be used to prevent flow separation.

Look at the diagram again, the combustion chamber mass flow rate is over 3200 lb/sec and the GG is around 215, not even 7%.
« Last Edit: 08/04/2016 02:52 AM by Jim »

Offline IainMcClatchie

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Re: Landing an MVac
« Reply #36 on: 08/04/2016 08:43 PM »
The proposal isn't to prevent flow separation, but rather to cause flow separation in an overexpanded nozzle to be stable, primarily to keep the unstable flow from smashing the nozzle.  There might be Isp advantages but I'm doing so badly already I don't want to get into that.

The main flow in the F-1 nozzle at the exit plane has to be a bit above 2600 m/s, because the sea-level Isp is 263 seconds.  The F-1 is overexpanded at sea level, so there is some pressure loss.  There's also a some lower-speed GG flow at the periphery of the nozzle.  Both these effects imply that the flow of gas from the main combustion chamber mas to be over 2600 m/s at the bell exit face.  So that's where that number came from.

I guessed 500 m/s for the GG exhaust.  The maximum theoretical velocity for that stream would be something like 1100 m/s.  But it doesn't go through a simple converging-diverging nozzle.  The more complex duct is going to lose energy, and as I said I was assuming it was injected subsonically.

Mixing increases with increasing shear, but it's a scale-dependent thing.  Hayward-Gordon has a somewhat useful brief introduction.

Pressure decreases from the sonically-choked throat to the exit because the flow is supersonic.  If the flow was subsonic and diverging, pressure would increase.  Obviously the boundary layer flow along the bell has to be subsonic, but I think the boundary layer has huge viscous losses that eliminate any pressure increase.

Offline Space Ghost 1962

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Re: Landing an MVac
« Reply #37 on: 08/04/2016 09:09 PM »
First - don't try to land (lower atmosphere) a US on a MVac. Not at all what its designed for.

Stay only in the "nonsensible" atmosphere. Because that's as low as the US can go, as a propulsion system.

Second - the majority of counter force / "thrust" and drag comes from stagnation pressure and engine operation.

This increases the density and size of the wake surrounding the stage, inflating it like a balloon. Which is torn apart in vortices, which carry away momentum, hopefully above the stage.

Extreme deceleration is the only point of this. You cannot get anything more then what's described here, and you're minimum cost is the props to decelerate with drag and thrust down to transonic. Then, it becomes a different problem.

In this environment, there is only having the drag of a huge wake, and the counter force of the momentum transferred to the combustion products which form the enormous, flattened, wake, which dissipates the total flux.

The transient of EI is where the mechanical effect, first only thermally, comes into play. One leverages the density increase against the flattened plume to transfer momentum through high velocity combustion products for maximum efficiency.

This isn't improved with a mechanism in any way. Only way it can get better is ... iSP. And more props for recovery. 

So stop messing with the engine.

Online drzerg

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Re: Landing an MVac
« Reply #38 on: 03/17/2017 04:37 AM »
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.   

Offline RotoSequence

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Re: Landing an MVac
« Reply #39 on: 03/17/2017 05:02 AM »
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.

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.
« Last Edit: 03/17/2017 05:04 AM by RotoSequence »

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