Author Topic: Rocket Engine Q&A  (Read 129032 times)

Offline kevin-rf

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Re: Rocket Engine Q&A
« Reply #480 on: 03/17/2013 03:20 AM »
You know what is more interesting? You started a thread on using a scud (vehicle derived from the Wasserfall) as the first stage of an orbital launchers. It's been done, Vanguard ...
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Offline Danderman

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Re: Rocket Engine Q&A
« Reply #481 on: 03/17/2013 05:00 AM »
You know what is more interesting? You started a thread on using a scud (vehicle derived from the Wasserfall) as the first stage of an orbital launchers. It's been done, Vanguard ...

As I am finding out.

But, Vanguard was a derivative of a derivative of a Scud ancestor (Wasserfall) and my issue in the other thread is whether an actual Scud could be used as a satellite launcher.
« Last Edit: 03/17/2013 05:02 AM by Danderman »

Online mmeijeri

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Re: Rocket Engine Q&A
« Reply #482 on: 03/22/2013 07:12 AM »
Terminology question: is the injector of the main combustion chamber considered part of the powerhead of an engine?
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Offline Antares

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Re: Rocket Engine Q&A
« Reply #483 on: 04/05/2013 06:55 PM »
No.  The powerhead is the turbopump(s) and gas generator or preburner.
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Offline kalif3000

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Re: Rocket Engine Q&A
« Reply #484 on: 06/04/2013 08:50 AM »
Hello everybody,

I am Faissal and new in the forum. I hope that is sufficient as intro :) .
I find this forum the best place for up to date info on current spaceflight topics, a really great source for news. I think many members here are experts in many aerospace related flields which is why I think I can finally find an answer to a (I think basic) question that is bothering me for a while here.

I am currenty studying the design of rocket engines and after scouting many books and other sources I hit a roadblock.
I noticed that all the equations I have seen so far that assess the performance of rocket engines use chamber pressure as input variable. However, those equations and the theory I have seen so far does not continue to elaborate on how the desired chamber pressure dictates the geometry and design of the rocket combustion chamber (or, the other way around, how to arrive at the chamber pressure from given chamber geometry and gas property)

I have seen some claimes here and there that the chamber pressure is equal to the pump discharge pressure minus the losses in the pipes (unfortunately not mentioned specifically in the books and good sources I have). Is that true?
If yes, then it seems counterintuitive to me that the size of the combustion chamber doesn't matter or am I missing something here?

I have recently whatched a short presentation where Tom Mueller explained how the Merlin engine works. There he explained that the turbopump discharges the propallent at 1400 and 1500 psi while the chamber pressure is 1000 psi. Where is this difference then coming from. It can't be all losses in the piping, or is it?

http://moonandback.com/2013/03/31/the-merlin-engine-presented-by-spacexs-tom-mueller/

Thank you for any advice and I apologize for any errors in the text. I am german, and not even a pro writing texts in my own language.


Faissal

Offline strangequark

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Re: Rocket Engine Q&A
« Reply #485 on: 06/04/2013 10:47 PM »
Hello everybody,

I am Faissal and new in the forum. I hope that is sufficient as intro :) .
I find this forum the best place for up to date info on current spaceflight topics, a really great source for news. I think many members here are experts in many aerospace related flields which is why I think I can finally find an answer to a (I think basic) question that is bothering me for a while here.

I am currenty studying the design of rocket engines and after scouting many books and other sources I hit a roadblock.
I noticed that all the equations I have seen so far that assess the performance of rocket engines use chamber pressure as input variable. However, those equations and the theory I have seen so far does not continue to elaborate on how the desired chamber pressure dictates the geometry and design of the rocket combustion chamber (or, the other way around, how to arrive at the chamber pressure from given chamber geometry and gas property)

I have seen some claimes here and there that the chamber pressure is equal to the pump discharge pressure minus the losses in the pipes (unfortunately not mentioned specifically in the books and good sources I have). Is that true?
If yes, then it seems counterintuitive to me that the size of the combustion chamber doesn't matter or am I missing something here?

I have recently whatched a short presentation where Tom Mueller explained how the Merlin engine works. There he explained that the turbopump discharges the propallent at 1400 and 1500 psi while the chamber pressure is 1000 psi. Where is this difference then coming from. It can't be all losses in the piping, or is it?

http://moonandback.com/2013/03/31/the-merlin-engine-presented-by-spacexs-tom-mueller/

Thank you for any advice and I apologize for any errors in the text. I am german, and not even a pro writing texts in my own language.


Faissal

Hi Faissal,

Chamber pressure is equal to pump discharge pressure minus all down stream losses, not just in "pipes" (propellant feed lines). The other big loss is across the injector. Injector pressure drop is usually 15-25%. Most of this pressure gets converted to propellant injection velocity, which helps break up propellant streams into drops and mix them well. It also keeps the chamber from "chugging". Merlin for example might have 5% loss in lines, then 25% through injector. 0.7*1500=1050psi.

"Chugging" happens when the injector pressure drop is too low. A small, random increase in chamber pressure will cause a large decrease in propellant flow. This will cause a large decrease in pressure, which causes a large increase in propellant flow, and so on in a cycle.

Pressure does not depend on chamber size and shape. Throat size is set directly by mass flow requirement. Chamber diameter is set as a multiple of throat diameter, based on experience with a given thrust. The designer must balance having a short convergent nozzle section, with keeping the chamber a reasonable diameter. You want neither a "pancake" or a "stovepipe", but there is some flexibility. Length is set to allow enough time for reaction, while keeping the chamber as light as possible. This is often set by defining a characteristic length (L*), which is Volume/Throat Area, and using a value, based on experience with a given propellant combination. It could also be set from first principles by looking at reaction time compared to residence time.

Chamber pressure is set to a desired value. Throat area, injector area, and pump output are selected for this value. You do not design the throat, injector, and pump and then calculate pressure, which is a misconception I remember having when I was in school.

Hope this helps, and welcome to the forum.
« Last Edit: 06/04/2013 10:48 PM by strangequark »
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Offline kalif3000

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Re: Rocket Engine Q&A
« Reply #486 on: 06/05/2013 07:41 AM »
Hi strangequark,

Yes, that was very helpful. I was planning to follow up on my question with one about the injector but you answered it all. Thanks!  :)

Offline Oli

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Re: Rocket Engine Q&A
« Reply #487 on: 07/27/2013 09:43 AM »

I have a question. Scaling up rocket engines seeems to be difficult/costly. I'm not talking about changing the cycle or increasing chamber pressure, only about increasing thrust. I wonder why this is the case and to what extent higher thrust engines are more expensive in development as well as in production.

Offline Jim

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Re: Rocket Engine Q&A
« Reply #488 on: 07/27/2013 09:46 AM »

I have a question. Scaling up rocket engines seeems to be difficult/costly. I'm not talking about changing the cycle or increasing chamber pressure, only about increasing thrust. I wonder why this is the case and to what extent higher thrust engines are more expensive in development as well as in production.

production and testing facilities have to be larger, hence most cost

Offline Oli

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Re: Rocket Engine Q&A
« Reply #489 on: 07/27/2013 10:05 AM »
^

Ok, so if you had to give development and production cost (seperate) as a function of thrust, how would it look like? Log, fractional, linear, quadratic, exponential?

Edit: Well I guess its not possible to describe it with a simple function.

Anyway, I've got another question  :)

Edit2: Is there a measure for engine efficiency independent of nozzle expansion ratio? Usually vacuum ISP for a given expansion ratio is given for an engine.

« Last Edit: 07/28/2013 10:23 AM by Oli »

Offline Proponent

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Re: Rocket Engine Q&A
« Reply #490 on: 07/29/2013 08:13 AM »
Well, there's c*, which is defined as chamber pressure times throat area divided by mass flow rate and is measured in velocity units.  When multiplied by the thrust coefficient, CF, a dimensionless factor defined as thrust divided by the product of chamber pressure and throat area, it gives the exhaust velocity: c = CFc*.

You can think of c* as a a measure of what happens in the combustion chamber and nozzle inlet and of CF as a measure of the efficiency of the expanding part of the nozzle.

Offline R7

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Re: Rocket Engine Q&A
« Reply #491 on: 07/29/2013 10:53 AM »
Edit2: Is there a measure for engine efficiency independent of nozzle expansion ratio? Usually vacuum ISP for a given expansion ratio is given for an engine.

IIRC you asked for theoretical efficiency with infinite expansion too before edited the question? Equation for that is

vexhaust = sqrt( 2*k*R*T/(k - 1) )

where k is ratio of specific heats, R specific gas constant (universal gas constant divided by average molecular weight of exhaust) and T combustion chamber temperature. Divide the result by g to get theoretical Isp in seconds. Note how result is independent of initial combustion chamber pressure and finite despite area ratio grown to infinite.

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Offline Oli

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Re: Rocket Engine Q&A
« Reply #492 on: 08/02/2013 12:58 PM »
Quote from: R7
Note how result is independent of initial combustion chamber pressure and finite despite area ratio grown to infinite.

Does that mean chamber pressure is irrelevant for vacuum ISP as long as P_e is zero? I.e. the bigger P_e in the term below, the more P_c matters.

1 - (P_e / P_c)^((k-1) / k)

So basically at sea level you want engines with high chamber pressure because P_e is high.

Consequently in vacuum higher chamber pressure makes only sense if it reduces nozzle size/mass (or is q somehow related to P_c?).

(variables from http://www.braeunig.us/space/propuls.htm)
« Last Edit: 08/02/2013 01:07 PM by Oli »

Offline R7

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Re: Rocket Engine Q&A
« Reply #493 on: 08/02/2013 03:32 PM »
Does that mean chamber pressure is irrelevant for vacuum ISP as long as P_e is zero? I.e. the bigger P_e in the term below, the more P_c matters.

1 - (P_e / P_c)^((k-1) / k)

So basically at sea level you want engines with high chamber pressure because P_e is high.

Yes, yes and yes. You want the P_e/P_c ratio to be as low as possible, zero P_e drives it to zero as long as P_c is non-zero to not to throw division by zero error.

In real world you have lower limit for P_e, even in vacuum, because temperatures get so low that your non-ideal exhaust gases stop behaving ideally and eventually condense.

Quote
Consequently in vacuum higher chamber pressure makes only sense if it reduces nozzle size/mass (or is q somehow related to P_c?).

(variables from http://www.braeunig.us/space/propuls.htm)

Things get quite unpractical if you drop P_c so much that required upper stage vacuum nozzle diameter exceeds that of LV's (like Sea Dragon concept).
mass flux q is directly related to P_c due to the good old PV = nRT. See equation 1.26 on your link (great page btw!). Equation solved for q(dot m) is;



(Sutton notation, t = throat, 1 = combustion chamber)
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Offline Oli

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Re: Rocket Engine Q&A
« Reply #494 on: 08/02/2013 03:56 PM »
Quote from: R7
mass flux q is directly related to P_c due to the good old PV = nRT. See equation 1.26 on your link (great page btw!). Equation solved for q(dot m) is;

So if you want to increase q, and thus thrust, you must increase A_t (and hence the size of the combustion chamber as well as the nozzle, I guess) and/or P_c.

Offline baldusi

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Re: Rocket Engine Q&A
« Reply #495 on: 08/02/2013 04:09 PM »
I've been thinking, about using different propellant for the TP of the GG (like the RD-107/8), but on the expander cycle case. If you were to do a bleed expander rocket, wouldn't using He to expand and move the TP give better performance? The Heat of vaporization of H2 is 0.904kJ·mol−1, while the He's is 0.0829 kJ·mol−1. If I understand that right, it would allow for a lot extra thrust for a given heat output. Basically, allow for expanding the thrust envelope of the expander cycle. I'm wondering if it would give better performance in the limited case. The tanking would be really bad, but the pressurization system would be "free". And if you direct the TP out gas to the nozzle the molecular weight should give some extra isp, right?

Offline Robotbeat

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Re: Rocket Engine Q&A
« Reply #496 on: 08/02/2013 11:23 PM »
Does that mean chamber pressure is irrelevant for vacuum ISP as long as P_e is zero?...
Yes...
Besides the impracticality of the bell becoming ridiculously big (which you mentioned), you also have the limit of condensing (or freezing) the exhaust as the temperature drops which limits performance.
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Offline R7

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Re: Rocket Engine Q&A
« Reply #497 on: 08/03/2013 09:26 AM »
So if you want to increase q, and thus thrust, you must increase A_t (and hence the size of the combustion chamber as well as the nozzle, I guess) and/or P_c.

Correct. Note that Isp is very non-linear function of area ratio so if engine diameter is restricted a quick'n'dirty solution is to just increase A_t (and CC) for larger thrust increase but minor Isp decrease. Attached example diagram for NK-33 (actual area ratio is 28).  The k value also affects thrust so propellant choice matters.

Besides the impracticality of the bell becoming ridiculously big (which you mentioned), you also have the limit of condensing (or freezing) the exhaust as the temperature drops which limits performance.

You appear to have skipped the next paragraph I wrote but no worries, the matter is important and deserves to be mentioned twice  :D
« Last Edit: 08/03/2013 09:27 AM by R7 »
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Offline Robotbeat

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Re: Rocket Engine Q&A
« Reply #498 on: 08/03/2013 04:24 PM »
Correct you are!  :)
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Offline Remes

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Re: Rocket Engine Q&A
« Reply #499 on: 08/04/2013 03:39 PM »
I've been thinking, about using different propellant for the TP of the GG (like the RD-107/8), but on the expander cycle case. If you were to do a bleed expander rocket, wouldn't using He to expand and move the TP give better performance? The Heat of vaporization of H2 is 0.904kJ·mol−1, while the He's is 0.0829 kJ·mol−1. If I understand that right, it would allow for a lot extra thrust for a given heat output. Basically, allow for expanding the thrust envelope of the expander cycle. I'm wondering if it would give better performance in the limited case. The tanking would be really bad, but the pressurization system would be "free". And if you direct the TP out gas to the nozzle the molecular weight should give some extra isp, right?

I think this doesn't make sense. You have to vaporize the fuel and oxydizer anyway. Be it in the thrust chamber or, with a expander cycle, in the nozzle channels. But the energy must be invested anyway.

If you take this energy now to heat up helium, you still have to vaporize Fuel/Oxidizer, so you didn't save anything. Furthermore you have a lot of hot helium which can't be used for anything (Edit: after it served its purpose to spin the TP), which doesn't provide thrust (as it can't be convinced to do any useful chemical reaction) and which was nevertheless heavy.
« Last Edit: 08/04/2013 03:52 PM by Remes »

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