Not forgetting SABRE engines on Skylon, of course. Well past TRL 3 by now, I believe.cheers, Martin
Say we have the money and charter to design and build a new engine for an SSTO space craft, (wings optional), with what design would we start and what technology would we incorporate? What would be the target Isp and realistic target mass ratio?
Quote from: aero on 06/26/2012 12:28 amSay we have the money and charter to design and build a new engine for an SSTO space craft, (wings optional), with what design would we start and what technology would we incorporate? What would be the target Isp and realistic target mass ratio?Not really going to happen with chemical propulsion and have a payload worth carrying.
For the purpose of this question thread, let's do forget about SABRE. It is the competition and the tech I am asking about is new rocket engine tech, not an air breather.
SABRE is a rocket. In air-breathing mode, it cools air and injects it in the chamber as the oxidizer. Once the air is gone, it injects LOX instead. So it is actually a rocket that uses an aerodynamic effect, rather than a jet engine that works with no air.And I should add that there are a lot of "cutting edge engine techs" that uses aerodynamic effects; aerospikes, ducted rockets, and expansion-defection nozzles to name a few. This is understandable, as such technologies could greatly reduce the cost of launch vehicles.
Jet propulsion is a means of locomotion whereby a reaction force is imparted to a device by the momentum of ejected matter.Rocket propulsion is a class of jet propulsion that produces thrust by ejecting stored matter, called the propellant.Duct propulsion is a class of jet propulsion and includes turbojets and ramjets; these engines are also commonly called air-breathing engines. Duct propulsion devices utilize mostly the surrounding medium as the "working fluid", together with some stored fuel.
Recent test results from the SABRE project can be found here;http://www.parabolicarc.com/2012/11/28/44703/#more-44703Sorry about stepping on your OP aero but this looks promising.
Not really going to happen with chemical propulsion and have a payload worth carrying.
Further detailed question:What exactly would be considered 'achievable' state of the art for a staged combustion hydrolox engine aiming at the same zone that SSME was aiming at 40 years ago? (Bleeding edge)SSME hit 3,000 psia combustion chamber pressure, so what would we be aiming at now?Assume for the purposes of this exercise, you can get SSME level funding annually once it ramps up ($700 million/year), and that leadership is tolerant of major catastrophic failures on the test stand, etc.
Thanks for the detailed reply, strangequark.A quick search on NTRS turned up this paper, which might be of some help to others:Advanced Oxygen-Hydrogen Rocket Engine Study by Aerojet; circa April 1981. (NAS 8-33452)(LINK to NTRS)The summary at the end was for hydrocarbon engines:The specific conclusions (based on the assumptions given in Table LVIII) derived from this study are as follows:(1) RP-1-cooled engines are limited to a chamber pressure of about 8960 kN/m2 (1300 psia) because of fuel coking in the coolant jacket at higher pressures.(2) RP-1-cooled engines with carbon deposit on the chamber walls are limited to a chamber pressure of about 13790 kN/m2 (2000 psia) because of coking of the fuel in the cooling jacket.(3) Refined RP-1 (e.g., JP-7) cooled engines are limited to a chamber pressure of about 17230 kN/m2 (2500 psia) because of specific Impulse (gas generator cycle) and to about 22060 kN/m2 (3200 psia) because of power limit (staged combustion cycle).(4) Refined RP-1-cooled gas generator cycle engines with carbon deposit on the chamber walls are specific Impulse limited to a chamber pressure of about 18610 kN/m2 (2700 psia).(5) LOX-cooled engines are specific Impulse limited to a chamber pressure of about 17230 kN/m2 (2500 psia) and 21370 kN/m2 (3100 psia), respectively, as gas generator and staged combustion cycles.(6) LOX-cooled gas generator cycle engines with carbon deposit on the chamber walls are specific impulse limited to a chamber pressure of about 18610 kN/m2 (2700 psia).(7) LCH4-cooled engines are specific impulse limited to a chamber pressure of about 20680 kN/m2 (3000 psia) and 24130 kN/m2 (3500 psia), respectively, as gas generator and staged combustion cycles.( LC3H8-cooled engines are specific impulse limited to a chamber pressure of about 20680 kN/m2 (3000 psia) and 24820 kN/m2 (3600 psia), respectively.(9) LH2-cooled LOX/HC gas generator cycle engines are power limited to a chamber pressure of about 37920 kN/m2 (5500 psia).
Thanks for the detailed reply, strangequark.A quick search on NTRS turned up this paper, which might be of some help to others:Advanced Oxygen-Hydrogen Rocket Engine Study by Aerojet; circa April 1981. (NAS 8-33452)(LINK to NTRS)The summary at the end was for hydrocarbon engines:(7) LCH4-cooled engines are specific impulse limited to a chamber pressure of about 20680 kN/m2 (3000 psia) and 24130 kN/m2 (3500 psia), respectively, as gas generator and staged combustion cycles.
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?
2) How hard would it be to increase the NK-33's chamber pressure/thrust while maintaining its thrust/weight ratio?
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,
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? -----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?
2) What are the major design difficulties in retractable engine nozzles?