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 ...
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
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.
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.
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)
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;
Quote from: Oli on 08/02/2013 12:58 PMDoes that mean chamber pressure is irrelevant for vacuum ISP as long as P_e is zero?...Yes...
Does that mean chamber pressure is irrelevant for vacuum ISP as long as P_e is zero?...
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.
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.
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?