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Cutting edge engine tech (Q&A) by aero on 26 Jun, 2012 00:28
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There are ideas for new tech scattered throughout this forum, and many appropriately over on advanced concepts. Many of them are mutually exclusive. My question is this:
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?
Please avoid nuclear, observe environmental and safety and stick with TRL-3 and higher tech. I guess I'm asking about the favorite future engine technologies on this forum. I know of two that interest me, full staged combustion and TAN. Others, not so much. -
#1 by MP99 on 26 Jun, 2012 11:50
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Not forgetting SABRE engines on Skylon, of course. Well past TRL 3 by now, I believe.
cheers, Martin -
#2 by aero on 26 Jun, 2012 14:50
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Not forgetting SABRE engines on Skylon, of course. Well past TRL 3 by now, I believe.
cheers, Martin
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. -
#3 by Jim on 26 Jun, 2012 14:57
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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?
Not really going to happen with chemical propulsion and have a payload worth carrying. -
#4 by aero on 26 Jun, 2012 15:08
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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?
Not really going to happen with chemical propulsion and have a payload worth carrying.
My hope is to invalidate your statement, Jim, but if we take our best shot and validate it, so be it. But let's take the shot. -
#5 by edkyle99 on 26 Jun, 2012 15:56
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For an SSTO, the rocket equation tells us that, to reach LEO, the following must be solved, roughly speaking.
e^(939/ISP) = Minitial/Mfinal
where ISP is the average specific impulse provided during the entire ascent (which means it must be something less than the vacuum ISP), and the right side of the equation is the mass ratio of the rocket and payload at the start, and end, of the propulsion system burn.
A rocket design is built around propulsion, so starting with an LH2/LOX system making an average 450 sec ISP requires a better than 8.045 mass ratio to make orbit. To put up a 9 tonne Falcon 9 v1.0 style payload, you would need a single stage rocket bigger than an entire Delta 4, that had a somewhat better dry mass fraction than a Delta 4, powered by something like three SSMEs, except that the engines would have to be even more efficient than SSMEs.
The mass fraction and ISP numbers are all within the realm of existing or nearly existing technology, but that's not the problem. The engines alone cost more than twice as much as a Falcon 9. Making the SSTO recoverable would be something else entirely.
The alternative would be to just buy an expendable Falcon 9, or Delta 4M, or Atlas 5.
- Ed Kyle -
#6 by simonbp on 26 Jun, 2012 16:32
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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. -
#7 by BobCarver on 27 Jun, 2012 00:15
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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.
Indeed, SABRE is a rocket engine (the RE is for "Rocket Engine") and has an ISP of 3600 in air-breathing mode. And, that basically blows the "rocket equation" out of the discussion.
But, I'm wondering what the limits of this discussion are. Jim said, "Not really going to happen with chemical propulsion and have a payload worth carrying." We can debate whether that statement is true or not, but there are propulsion concepts which do not depend on chemical reactions (or nuclear), such as metallic hydrogen or MHD, so are we limiting this discussion to traditional rocket engines which carry all of their oxidizer and only use chemical reactions to generate thrust? -
#8 by QuantumG on 27 Jun, 2012 00:23
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From Sutton:Quote
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.
The SABRE design is a rocket/duct propulsion hybrid that can act in either mode.
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#9 by D_Dom on 29 Nov, 2012 22:29
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Recent test results from the SABRE project can be found here;
http://www.parabolicarc.com/2012/11/28/44703/#more-44703
Sorry about stepping on your OP aero but this looks promising. -
#10 by RanulfC on 30 Nov, 2012 00:50
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Recent test results from the SABRE project can be found here;
Uhm the "question" I guess would be ya, while this "looks promising" it looked JUST as promising when the Japanese did it 5-10 years ago. THAT didn't get off the ground either, so what makes this any "more" promising with less money and a longer time-scale for development?
http://www.parabolicarc.com/2012/11/28/44703/#more-44703
Sorry about stepping on your OP aero but this looks promising.
Aero: Plain and simple for SSTO you are PROBABLY going to have to "cheat" at some point, you are definitely going to have to if you insist on sticking with chemical rockets. It's quite realistic to get a SSTO rocket vehicle to work, you just have to ask who's going to allow a HEDM/FLOX SSTO to take off from a Spaceport near them
Randy -
#11 by RyanC on 01 Dec, 2012 13:20
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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?
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. -
#12 by mmeijeri on 01 Dec, 2012 13:32
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Not really going to happen with chemical propulsion and have a payload worth carrying.
Is your objection based on the absolute size of the payload or on specific launch costs? -
#13 by RyanC on 02 Dec, 2012 03:16
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Because nobody answered, here are some attempts to give answers:
In 1990, there was a 12-month study of currently undeveloped rocket engine cycles with applications for advanced transportation (NAS8-36643) by SRS Technologies for MSFC under "Manned Mars Mission and Program Analysis".
So what did they look at?
Split Expander (SX) Cycle
Basically, it's a derivitation of the conventional expander cycle, to attain engine thrust levels of 500 klbf, while still having a chamber pressure much less (<1,500 psia) than gas generator or staged combustion.
Full Flow Staged Combustion (FFSC) Cycle
Basically, all propellant flow is ultimately burned in the main chamber and exhausted through the nozzle, giving this cycle maximum ISP. Since all the propellant flow is being used to drive the turbines, it results in the lowest possible Turbine Inlet Temperature (TIT) for a given main chamber pressure.
Conceptual design studies showed that it offered the possibility of simplicity and low cost compared to conventional staged combustion, without compromising on safety/reliability/weight/performance compared to SSME-style staged combustion cycle.
End results gave these numbers for a SSME replacement (see attached)
The full document is at NTRS HERE.
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#14 by strangequark on 02 Dec, 2012 03:48
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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.
Well, there's an upper limit in what you can do, based on the turbomachinery power balance, and the maximum temperature you can allow in your turbine. For a given propellant, amount of required pump power is linearly proportional to the pressure rise. 3000psi is pretty close to the upper practical for hydrogen fuel-rich staged.
Delivered turbine power for 100% efficient turbine is:
mdot*Cp*Tin*(1-PR^((1-gamma)/(gamma)))
mdot=Mass Flow Through Turbine
Cp=Turbine Drive Gas Specific Heat
Tin=Turbine Inlet Temperature
PR=Turbine Inlet Pressure/Turbine Outlet Pressure
Gamma=Turbine Drive Gas Specific Heat Ratio
The long and short of all this is that to increase turbine power, and thus chamber pressure, you need to increase mass flow or turbine inlet temp. For a staged engine, PR can't be pushed very much, and Cp and gamma are just what your propellants give you. So mdot and Tin are the dials that can be easily twisted.
So, your turbine materials set how high your turbine temp can be. I think the SSMEs are around 1000-1100K. With enough money, you could practically get up to 1500K maybe. The downside is that your Cp goes down for fuel-rich drive gas as you add more oxygen to increase the temperature.
Full flow staged looks at the mdot part of the equation, by bringing the oxygen mass flow to bear. There's a lot of oxygen available in a hydrogen engine, but ox-rich gas isn't nearly as good as hydrogen rich gas as a turbine drive gas. The Cp on ox-rich gas is roughly a tenth that of h-rich gas. So, even though you have about 5 times as much oxygen, there's, very roughly, only 50% more turbine power to be added. Then though, you need more pump power to pump all of your oxygen up to preburner pressure, instead of just main combustion chamber pressure.
The bottom line is that hydrogen tops out at maybe 4000psi if you implement one of these, and very optimistically, maybe 5000psi for both.
Methane, using the same parameters, tops out at 8000-10000psi. Full flow staged combustion methane is my poison of choice if you're crazy enough to attempt an all rocket SSTO. That's an engine that is in the 350s at sea level and the 380s for vacuum, with 2.5 times the mean density of hydrolox. It's also a better fuel choice if you've really lost your mind and want to do reusable all-rocket SSTO. -
#15 by RyanC on 02 Dec, 2012 04:12
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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). -
#16 by strangequark on 03 Dec, 2012 14:04
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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).
Caveat to the above. It looks like they chose an arbitrary limit of 8000psi pump discharge pressure, which for the staged cycles will naturally limit you to about 3000-4000psi range on chamber pressure. They're leaving available power on the table to keep it under that limit. -
#17 by Hyperion5 on 31 Dec, 2012 06:13
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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.
Speaking of Aerojet and metholox engines, I've got quite a few questions involving them. The first set of questions involves the relatively recent proposals by Aerojet to build NK-33 derived advanced kerolox staged combustion engines.
I was once told by an Aerojet propulsion engineer that the basic design of the NK-33 was nowhere near being maxed out in terms of thrust & chamber pressure. He said it should thus be relatively easy to maintain the thrust/weight ratio while increasing the thrust from 338,000 lbf (SL) to 500,000 lbf (SL) via more propellant mass & increased chamber pressure/better Isp. This, from what I'm told, led Aerojet to commit itself to proposing an "all-new engine" called the AJ-1-E6 for NASA's SLS derived from the NK-33 but dual-chambered like an RD-180.
However I've never seen another propulsion engineer vet all that. That brings me to my first set of questions.
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 next set of questions concerns both the NK-33 and metholox engines. A few weeks back I was impressed with Strangequark's talking up the advantages of a staged combustion metholox engine being "beautifully simple" and high performance. It inspired me to try simulating a rocket with SC metholox engines with modemeagle's help. We wanted to see just how impressive Strangequark's engine setup would be, and even recently using more conservative figures, we have not been disappointed.
Although we now have a fairly good idea of what the engine and rocket performance should be, I'm not sure about our engine architecture. The rocket design we're using relies on a lot of SC metholox engines, 5 large (400,000 lbf) engines on SI, and 5 small (50,000 lbf) engines on the SII, for a total of 10 engines produced per expendable launch. Both sets of engines are essentially cousins of each other and share the same chamber pressure (2900 Psi) and design architecture. I also had the hope of eventually making both stages reusable via a Grashopper-like program.
Our problem is, we need a large number of relatively high-performance metholox engines that are fairly elegant designs, share as much design architecture as possible, can be produced quickly and in large quantities (especially the big engine), and would be robust enough to eventually handle multiple flights with relatively quick maintenance & refurbishment. My first choice of engine architecture was the NK-33's, which was to have been produced in massive quantities for the N-1F, has excellent performance, is a relatively simple staged combustion engine, is fairly robust and supposedly the design's layout can handle higher pressures than 2103 Psi. 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?
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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?
2) What are the major design difficulties in retractable engine nozzles?
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#18 by Nomadd on 31 Dec, 2012 11:43
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Has anything happened technology wise in the last few decades that could make the X-33 type aerospike worth revisiting? I know that engine mass and performance not coming in where they'd hoped was one the reasons it was cancelled, but they must have thought it was possible when they started.
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#19 by Damon Hill on 03 Jan, 2013 02:06
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I ran across a highly detailed document about the RD0120 (SSMEski) that showed a modified version with an expendable internal carbon-carbon 'choke' nozzle that allowed a fully-expanded main nozzle to be used. The choke was jettisoned at altitude. Kind of goes against the grain on keeping designs simple, but I'll throw that concept out on the table for discussion.
http://www.lpre.de/kbkha/RD-0120/index.htm
http://www.lpre.de/sntk/NK-33/index.htm
This site contains detailed documentation on a wide variety of Russian rocket engines and their development history.
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.