1) How hard is it to increase the thrust & chamber pressure on a staged combustion engine like the NK-33? What kind of limits are there to these increases if you strengthen the basic design?
Which chamber? You increase the main chamber pressure, you have to increase the pre-burner chamber pressure to compensate.
2) How hard would it be to increase the NK-33's chamber pressure/thrust while maintaining its thrust/weight ratio?
Well according to your friend not very. If the chambers can handle the pressure then it's a question of what you do to increase the pump flows. That depends how close to maximum capacity they already are. If they have plenty of margin it could be something as simple as reducing the size of a blanking plate on the inlets. If not then you're looking at new pump and turbine design.
3) What are the advantages and disadvantages of an NK-33's staged combustion system compared to that of the RD-191's?
The only modifications to the overall internal design architecture would be to run the pre-burner fuel-rich,
Why? If you want reusability you want a design where seal failure is not a criticality 1 condition. If you insist on on SC with the NK33 architecture that means you run Ox rich in the preburner so if the seal to the LOX pump leaks you're not putting a nice hot fuel stream in contact with a nice dense oxidizer. Making Sc engines more failure tolerant is why you'd run one PB fuel rich and the other ox rich and drive their respective pumps from those burners. The results of that Aerojet study suggest for easier maintenance you run with LOX cooling.
modify the injectors and plumbing to handle methane, and strengthen the design architecture to handle higher pressures.
However I've no idea how ideal an engine that looks like an NK-33 internally is for methane combustion or potential reusability. I've got a ton of design questions.
1) What design architecture would you recommend for these mass-produced metholox engines? Should we stick with something like a modified NK-33's design or move towards something else like the RD-191 or a full-flow staged combustion engine?
2) If you recommend a different design architecture, how would that affect performance and cost of production?
3) How should we cool these engines?
4) Can you make SC metholox engines with 2900 Psi chamber pressures relatively quick (say 1-2 weeks tops) to refurbish and fly again?
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I've got one last set of questions and it concerns retractable engine nozzles. To make the SII potentially reusable, it'll need to have retractable engine nozzles, as the 300+ expansion ratios the smaller engines feature will cause big problems at sea level.
1) How high of an expansion ratio could metholox engines tolerate at sea level if they're being used to land a stage?
Well the SSME took off from SL with an expansion ratio of 77. Rule of thumb in US is down to 0.4 of ambient pressure you should be safe from flow separation. Others reckon you can go to 0.34 of Pambient.
2) What are the major design difficulties in retractable engine nozzles?
Ask PWR they do a retracting nozzle version of the RL10. Needs 3 electric actuators and is RCC. Other concepts included the Bell Aerospace "rubber" nozzle using flexible foils with exterior ribs that extended, kind of like a sock.
Note if the engine is not firing it's
relatively simple to do. I'd presume you'll have plenty of time on orbit after you fire you're de-orbit burn to retract it to the Earth safe expansion ratio from the high Isp ratio.
I'm sticking to the thread title not the later stated goal of an SSTO.
I will note that the engine is indeed the critical element of enabling an SSTO.
And let me suggest people Google the sci.space.tech newsgroup for the last 20 years of posts on this subject.