Turborocket Upgrades
Jul 5th, 2009 by johnhare
guest blogger john hare
John Bossard over at Plasma Wind http://plasmawind.typepad.com/ introduced me to air turborockets after a propulsion conference in 2000. My hard references on the subject were from his recommendations. He has built operating engines on this cycle. Never being willing to let well enough alone, I suggest a few upgrades. He may take the ideas apart if he has time and interest, and it doesn’t conflict with his business.
A normal turborocket is an air breathing engine with a rocket gas generator driving a turbine which drives the air compressor. The fuel rich turbine exhaust mixes with the compressed air to burn in the afterburner. The result is an engine with over twice the thrust to weight of a jet engine with immunity to flameouts. It can reach higher altitudes and higher airspeeds than any turbojet, and is simpler to operate. The down side is that it sucks down far more fuel than a jet, and won’t operate in vacuum like a rocket.
The turborocket is a niche engine for air breathing acceleration from mach zero to mach five or so. It can actually do the things people try to claim a ramjet can do. It is also good for relatively short duration cruise when engine mass is very critical and the good thrust to weight ratio outweighs the high fuel consumption. It also uses far less fuel than a rocket when in its’ proper working environment.
A rocket may have an Isp (Dense Fuels) of 300 with a T/W of 100. A turbojet may have an Isp of 3,000 with a T/W of 10. A turborocket might have an Isp (dense fuels) of 750 with a T/W of 20. Clearly the turborocket needs either a better Isp, or better T/W to make it a clear win over the more traditional engine cycles. For this particular cycle, hydrogen makes a lot of sense.
Hydrogen gives a 30-40% Isp boost to rockets, at the expense of tripling propellant tankage per unit of propellant mass, when H2/O2 is compared to Kero/O2. It is an ongoing argument whether LH2 is good for launch vehicles. Upper stages benefit far more from the extra performance. Hydrogen on air breathing engines though, exhibit a 300% increase in Isp.
Hydrogen has about 3 times kerosenes energy per pound when burned with atmospheric air, so even being 10 times as bulky per pound, it only has tanks a bit over 3 times the size of the kerosene per BTU. Since acceleration engines have to lift the take off mass, cutting your fuel load by two thirds has to have an effect on engine mass, along with wing and landing gear mass of course.
Hydrogen has another valuable trait in its’ specific heat characteristics. It is the best possible coolant with about 16 times the cooling capacity of air. 16 pounds of air introduced to 1 pound of colder hydrogen will meet it about halfway on temperature. Hydrogen can cool enormous quantities of air either through a physical heat exchanger, or with Mass Injection Pre Cooling (MIPC). Many studies have used precooling in some form, though most of them just get too complicated for the simple functions we need for a launch system.
I suspect a straight switch to hydrogen from dense fuels will less than double Isp, considering the oxydizer mass to be carried. I suggest using the cooling properties of liquid hydrogen to precool a core flow for a turborocket. This potentially increases mass flow through the gas generator while much decreasing the on board oxydizer requirements.
After the initial compression of air, about a sixth of the total flow is sucked into the core for gas generator use. A pound of LH2 as injection precooling for each sixteen pounds of air per second will almost double the density compared to uncooled air that has been through the same compressor. True gas generators operating in ejector mode compress the air/hydrogen mixture and burn it in the middle chamber that drives the turbine. The combination of precooling and ejector compression should give a pressure in the middle chamber of about four times that of the main chamber. With a turbine pressure ratio of four, and a core mass flow of 20% of total compression, a fairly high main compressor ratio (for turborockets) should be achievable.
The turbine can be cooled by the remaining hydrogen required to reach stoichiometric burn in the main chamber. I suspect an Isp in the 2,000 range can be reached during acceleration periods, with cruise phase considerably higher. This while retaining most of the strengths of the turborocket system.
John,
great post, and thank you for mentioning plasma wind. As usual, you present a wealth of ideas for airbreathing propulsion concepts. I also appreciate your generosity in sharing your ideas. If it’s useful, I thought I might provide some additional comments on your post in the interest of promoting TBCC concepts in general, and the ATR concept, specifically.
As you mention, the ATR is an accelerator engine, and is really the only single airbreathing engine that can span the zero to mach 5-6 flight speed range range. This makes it a good choice as an endo-atomospheric intereceptor engine, and as an airbreathing component to a space launch system. Interestingly enough and almost as an aside, I would note that, in principle the GG-cycle ATR will in fact be able to produce thrust even in a vacuum, although at a far-off-design operating condition. The GG is going to keep running, because it has its own fuel and oxidizer. The thrust would come from whatever gas generator residual overpressure would result in the ATR’s combustor (or afterburner).
Typical Isp’s for ATRs range on the very low end of around 650 to 850 for solid propellant GG ATRs. Hydrogen GG ATRs can easily exceed 4000 seconds, with RP1/LO2 ATRs in the 1500 to 2500 sec range, and even higher for other hydrocarbon fuels like methane.
A hydrogen regen cycle ATR was in fact the cycle decided upon for use in project Suntan, a high-speed (mach 3+) high altitude interceptor concept developed in the 1950’s. Several engines were built and tested, and conclusively demonstrated it’s performance. However, the bulk of LH2 made the aircraft’s fuel tanks far too large, and the project was abandoned. Nevertheless, the technology development efforts required for this engine allowed Pratt and Whitney to learn how to generate, store, and handle cryogenic hydrogen. This knowledge underpinned and led to their subsequent development of all of their liquid hydrogen engine systems, including the phenomenally successful RL-10 engine.
As you know, I’ve had a long-term interest in the use of inlet-spray cooling for the enhancement of airbreathing engine performance. In recent years, this technology has been dubbed MIPCC for Mass-Injection for Pre-Cooling Compressors. The increasing stagnation temperature of the incoming airstream can be the limiting factor for airbreathing engine flight speed, so inlet spray cooling may be good way to transcend this limitation. However, you generally inject an oxidizer, or water, into the incoming airstream, and not a fuel. If you inject a fuel, especially hydrogen, the fuel can actually auto-ignite with the incoming airstream. This would be bad. Also, you need to inject in front of (upstream) of the compressor, rather than behind (downstream). Otherwise you don’t provide pre-cooling to the compressor, and you have to supply the coolant at a higher pressure.
In my opinion, I don’t think LH2 makes for a good fuel choice for the ATR. Despite its high performance numbers, it’s bulk density simply make for large fuel tanks. This seemed to have been the conclusion from project Suntan. However, maybe there’s some ways around this issue, so I’m willing to be convinced by a good idea and numbers, should they show up.
I’ve done quite a bit of work with liquid oxygen and with the fuels RP1 and propane. I like LO2 because it’s cheap, dense, and you can make it any place you’ve got electricity and air (which is pretty much everywhere on the earth). The kerosene group (the RP’s, the JP’s, the K’s) are also quite dense, and economical, and widely-available. However, I think a fuel that may offer a lot of promise for ATRs is liquid methane. It sort of has a density between hydrogen and kerosene, has high performance, and is widely available (any place where there’s cows). Like hydrogen, it has excellent cooling properties. In short, it’s cheap, somewhat dense, and widely available. I’ve only got my performance calculations speaking for methane ATR performance, but I’m preparing some new ATR hardware that I may be able to explore methane performance with in the near future.
This sort of became a long post, but I thought your post deserved some good commentary, so I hope this qualifies! Keep the posts coming, and we’ll continue the dialog here and at Plasma Wind. I’m prepping the next ATR post, which will address good ATR applications
John B.,
Good points. I didn’t realize Suntan was an ATR. BTW, have you looked at subcooled propane as a fuel for an ATR? Liquid propane is just barely still liquid at the normal boiling point of LOX, has a density at that temperature around that of kerosene, and has a performance intermediate between methane and kerosene. It’s also a pretty darned good thermal fluid. If you need LOX anyway for the precooling and gas generator, using subcooled propane as the fuel might make sense.
~Jon Goff
Just a quick note before “real-life” overwhelms me again…
)
Jon: Beat me to it, I was going to point out Cryo-SubChilled propane myself. Another “point-in-its-favor” is if you have LOX, you can run commercial (check your local propane specs though, a LOT of what is “called” propane isn’t as noted in the yarchives web-notes on the subject
) propane through the LOX to chill-out a lot of the “stenchents” and added chemicals which makes it a lot cleaner to run especially if you’re using a regenative cooling loop.
John H. and John B:
Having read up some of the articles and such on ATRs what about using your “cryo” component (fuel or oxidizer, but specifically I’m thinking Liquid Propane, Liquid Methane, or Liquid Hydrogen) run through a heat exchanger or through a regenerative cooling system to “run” your turbine and then mixing it with the compressed air and thence your combustion chamber? Is there any specific reason that an ATR HAS to run a “hot” gas generator?
I’ve seen patents and articles on engines that are for all intents and purposes “ATR” engines with the compressors connected to turbines spun by expanding cryo-fuels (usually LH2) some going as far as making the ‘gas-generator’ portion of the engine an actual rocket engine deep-throttled to a very low setting to super-heat the fuel before it get to the turbine.
(Some will have the rocket combine the compressed air with the fuel and then burn it to produce the thrust, some even going as far as wrapping a ramjet duct around the engine and injecting MORE fuel down-stream)
Thanks for the information.
Randy
Information overload, I’ll reply when I have time to digest some of it.
Jon Goff:
in answer to your question: no, I haven’t looked at sub-cooled propane. But your suggestion makes a lot of sense. That would certainly improve impulse density performance. I will certainly have to look into that in greater detail. In running my ATR with propane, we used it an ambient temperature allowing us some autogenous pressurization for modest gg pressures.
Because methane and propane have intermediate densities as well as Isp performance, they may represent a sort of “optimum” between hydrogen and the kerosenes. Its the large amount of hydrogen in methane and propane that allow for their excellent thermal capacity properties. One thing you have to be careful of with methane and propane is coking within regen passages at elevated temperatures, and also carbon formation during combustion at fuel rich conditions within gas generators. When I was at Aerojet, we did a good deal of work on both of these issues, mapping out temperatures and mixture ratios that could avoid problem areas. Anyway, thanks for the suggestion.
Randy:
you make an important point regarding additives to liquid hydrocarbon fuels. We found that certain compounds, specifcially any sulfides such as might be found in LNG (LNG: liquified natural gas, comes from wells, ususally around 93% or better methane) could induce the more rapid onset of coking. “stenchents’ may certainly have these infuences as well.
The idea of running a LH2, LCH4, or LC3H8 through a regen loop and then using the heated fuel to drive a turbine is a excellent idea. So good in fact, that it is the very engine cycle used for project Suntan: Behold, the model 304, an LH2 regen-cycle ATR.
The best reference on project Suntan and the amazing model 304 ATR can be found in “Liquid Hydrogen as a Propulsion Fuel: 1945-1959″, by John L. Sloop, a NASA History series publication, NASA SP-4404, Chapter 8.
Regarding your question about the requirement for a hot GG: No, there is no fundamental requirement that ATRs must have a hot GG. The requirement is that the turbine drive gas have large amounts of available energy that can be extracted efficiently. However, there are fundamental differences in GG-cycle ATRs compared to regen-cycle ATRs. A more lengthy discussion would be required to elaborate on these differences, but perhaps more of the more important differences is: the regen-cycle produces a higher Isp, but a lower thrust-density compared to the GG-cycle.
The deciding question for determining whether or not a TBCC is an ATR or not is this: Does any atmospheric air touch the turbine? If it does, its not an ATR. If it doesn’t, it can be counted as an ATR. Most all ATRs “afterburn”, which is to say that the turbine drive gas is fuel-rich (i.e rich combustion in the Gas Generator, for example) and is subsequently burned with air from the compressor in a combustion chamber. This chamber is essentially an afterburner in turbojet parlance, but for ATRs the term “combustor” is the accepted nomenclature. There are a few ATR designs however, that do not “afterburn”. In my opinion, the most notable of all these non-afterburning ATR designs comes from none other than…..Robert H. Goddard. In the March 1932 issue of Scientific American, he published an article describing a new propulsion system, for improving the performance of aircraft propusion that could fly much higher and faster, and that machine is essentially an ATR. Early prototypes of his rocket-driven fan engines can be found in the Goddard Museum in (no kidding) Roswell NW.
Well, again I’ve waxed on a bit here. If there is interest in further discussion, I’ve been writing a multi-part series on the ATR cycle on my blog: http://plasmawind.typepad.com/ .
I would look forward to further discussion.
Jon Goff:
In answer to your question: No, I have not looked at sub-cooled propane. But that’s a good suggestion. It would definitely help density Isp performance. That’s probably something I should really look into. To date, I have run ambient temp propane, mainly because it I can use the autogenous pressurization to obtain modest GG pressures.
Methane and Propane do indeed have excellent thermal capacity properties. These hydrocarbon fuels get their excellent heat capacity properties largely from their abundant hydrogen content. I believe that methane, and perhaps propane, may offer some sort of “optimum” performance between Isp and bulk density. Plus, as previously mentioned, that are readily available and economical.
Randy Campbell:
You’re quite correct regarding additives and impurities in the hydrocarbon fuels. They can really cause some problems and reduce performance, and for a variety of different reasons.
When I was at Aerojet, we did a lot of work with methane, RP1, propane, and LNG (Liquefied Natural Gas, which is usually 93% or better methane). With the use of hydrocarbons in regen flow passages, you can have coking, which forms hard coatings on the inside walls. In can reduce heat transfer enough so that you can get burn-throughs. The longer the hydrocarbon chain, the greater the susceptibility to coking. Methane and propane were not particularly susceptible to the coking, but under the right conditions, would coke. Another issue is solid carbon formation in combustion processes when running fuel rich, such as in fuel-rich gas generators. Methane never had a carbon deposition problem over the mixture ratios that we ran at. Propane only had a problem a rather rich conditions, but when it formed solid carbon, it was very hard and difficult to remove. RP1 formed carbon below O/F mixture ratios of about 0.35 and below.
There is no specific requirement that an ATR has to run a “hot” gas generator. The requirement is a source of high-enthalpy drive gas in which the energy can be extracted efficiently. Thermodynamically , this usually means a hot, low molecular weight gas driving a turbine. How you achieve that is left to the imagination of the inventor. Running LH2 through a heat exchanger and using that to drive a turbine is an excellent idea. So excellent, in fact, that this is the very configuration of the engine that powered the Suntan aircraft: Behold the Model 304, an LH2 regen ATR!
Indeed, there are many configurations that qualify as an ATR cycle. The criterion that I use for deciding whether an engine is an ATR or not is this: does any atmospheric air touch the turbine? If it does, it’s not an ATR. If it doesn’t, it may be an ATR-cycle.
The best description of project Suntan and the amazing model 304 engine can be found in “Liquid Hydrogen as a Propulsion Fuel, 1945-1959”, by John L. Sloop, The NASA History Series, NASA SP-4404, Chapter 8. It also covers similar ATR-like engine cycles such as the REX.
The ATR cycle is not particularly new. The first example that I know of for an ATR-cycle comes from none other that Robert H. Goddard. In 1932, Dr. Goddard published an article describing what is basically an ATR engine for powering aircraft to high altitudes and speeds. Prototypes of his GG-powered fans, which are essentially ATR engines, can be found in the Goddard Museum in (no kidding) Roswell, New Mexico.
Again, I’ve waxed on a bit here. I’ve been writing a multi-part discussion of the ATR at my blog “Plasma Wind”, which I think may have some useful discussion there, and I would welcome further discussion
This is a fascinating post and thread. I have nothing to add technically, as it’s beyond my knowledge level to do so. I do have some anecdotal, historical comments on ATRs to add, FWIW. Dick Morrison was a consultant for Aerojet back in the 80’s (he headed the team in the late 50’s that designed and built the Thor ORBM). We had a number of conversations about ATRs and their possibilities. He actually considered ATRs as viable strap-on boosters for vertical launch. He referred to the ATRs as “cans,” and pointed out that Aerojet had built a 6 ft. diameter ATR back in the 60’s.
ATRs are another one of those terribly under-used propulsion concepts that the aerospace policy process has left littering the procurement graveyard. Too bad, really, and I hope that changes soon.
Morrison also came up with a concept to extend the Mach range of ramjets: the Oblique Detonation Wave Engine, or some such — I remember the acronym as being “ODWE.” The idea was to enable combustion at higher air inlet speeds without moving parts to change the inlet configuration.
http://www.google.com/search?hl=en&rlz=1T4GGLG_enUS312US312&q=oblique+detonation+wave+engine&aq=0&oq=Oblique+Detonation+Wave&aqi=g1
The ODWE simply added a “speed bump” to the “floor” of the engine to affect the shock wave pattern. The idea was that you could Mach 7-8 speeds with it. I informally lobbied Congress to consider it as an alternative concept to the more ambitious NASP technologies, but to no avail.
John B.:
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Yep, my “opps” I finished reading your blog AFTER I posted
(Teach me… well, no.. not really
The ‘trick’ with Propane is to use “pure” propane as far as I’ve been able to read. Most of the stuff was discussion on alternate fuels run through the RL-10 at various times during testing and involved the coking properties of various types of hydrocarbon alternatives to straight LH2 fuel.
)
Question: Does one have to set up an account to access the “discussion” portion of your blog? I can access the blog itself from my work computer but can’t log in it seems.
(Yes, yes, I know… the horror of a government employee who has time to read stuff on the internet at work! Lets just say there are days when my job doesn’t suck as much
Randy
How would the cost of an air turborocket compare to a rocket of equal thrust?
The discussion in “The Rocket Company” mentioned the idea of using airbreathing engines on their pop-up first stage to extend the hover time during vertical landing over what was possible with rockets of the same mass. They seemed to think that the cost of air breathing engines was cost prohibitive for any part of the trajectory other than landing the stage almost empty.
And yet, the zero to mach 5 speed range is almost exactly what you would need for the first stage of a pop-up trajectory. Could you economically power such a stage with “cans” of this type for the first 100,000 feet and first 2000m/s?
The concept of “strap-on” or “can(s)” boosters for a rocket has come up before, though not here as I recall. I found a lot of info on the internet and yarchive using the search terms “jet-booster first stage” or “air breathing launch assist” or “first stage jet engines” in various search engines.
Danni Edar(sp?) was quite prolific an one point with the idea and pointed to work done for Boeing on such a concept for a low-captial investment launch vehicle using Mil-Surplus F-100 low-bypass turbofans (ex-F-15, current F-16 jet engine) for around the first 60 seconds of launch, and a three barrel (two-stage) vehicle powered by 6 (2 on each barrel) RL-10 rockets.
(All engines were started on launch)
Due to the air breathing engines the angle of launch went from 90 degrees at “launch” to around 35 degrees (@Mach-2 at 50,000ft) at the point where the ten (10) jet engines were ’staged’ and the vehicle continued on to orbit with @6,000lbs of payload. All parts were recovered for re-use.
Have to look up some of the stuff, (I’m moving at the moment so that may take a bit) but how much extra thrust would an ATR “booster” provide and would it all be worth the effort in doing all the needed aerodynamic and structural work to add to the vehicle to reach a higher Mach number before staging?
As an added ‘note’ here: Any acceleration done by a vehicle over Mach-1 above 100,000ft is supposed to NOT cause audible sonic booms on the ground. Can an ATR operate that high?
Randy
Roderick Reilly:
I didn’t know Dick Morrison when I worked at Aerojet (1986-1990), but I would have liked to have met him. I actually went to work at Aerojet specifically to work on the ATR. We were developing a solid-propellant GG ATR, for missile applications.
Aerojet did indeed build and test a mono-propellant ATR in the 1960’s. I don’t think it was 6 ft in diameter, more like around 2 ft. It used catalytically-decomposed ethylene oxide (if I recall correctly). The modern concept of the ATR was actually developed at Aerojet in the 1950’s by William House. Many early ATR patents bear his name.
Randy Campbell:
I regret that you have been unable to access the “discussion” portion. (actually, I didn’t even know I had a discussion portion ). Try clicking on the “comments’ at the bottom of the relent post to provide some comments. I’ll try and track down what’s going on, and let you know. (also you can try contacting me via bossrd.com )
The ATR has indeed been proposed as a strap-on type booster. This may be a viable configuration. You couldn’t use turbojets in this fashion because their thrust-to-weight isn’t quite high enough, and ramjets and scramjets don’t make static thrust. For this application, the ATR is really the only viable airbreathing engine that would work. It is the ATR’s intrinsic simplicity that allows it be much like a “can” strap on. The April 1959 issue of Astronautics has an interesting article on using the ATR as a first stage booster for a Vanguard-sized small satellite launcher. Considerable reductions in overall gross lift off weight were realized by using an ATR instead of a liquid rocket (22,000 lbm for Vanguard vs. 11,300 lbm for the ATR-powered first stage).
And yes, our studies showed that an ATR could reach altitudes in excess of 120,000 ft at supersonic flight speeds. At a given flight speed, the ATR can (theoretically) fly at higher altitudes than any other airbreathing engine. This is because it retains the highest pressure in its combustor (for a given compressor ratio, or inlet compression ratio for ram/scramjets) compared to any other airbreather (see plasma wind blog post, part 1)
Jsuros:
In the broadest terms, with no stated assumptions, an ATR would cost somewhat more than a rocket of equal thrust, but not orders of magnitude more. This cost would come largely from the required turbomachinery: compressor and turbine. The ATR’s turbomachinery can be much simplier than that of a turbojet, so this makes for a more economic airbreathing engine.
I believe that you could use ATRs for the applications that you reference. Technically, such an application appears feasible. More analysis would have to be performed to determine if such a configuration would be economical, however.
What am I missing here? The ATR is an airbreathing engine that can lift you from the ground to a vertical trajectory at 120,000 feet and 1700m/s with an Isp in the 1000 range and T/W around 20. This technology has been known since the 1950’s, and yet no one has thought to use this for a suborbital vehicle?
This technology has been known since the 1950’s, and yet no one has thought to use this for a suborbital vehicle?
While I have thought of it, an Italian gentleman prevented me from acting on it. I believe you have met him. –Mr. Myfundsalow.
A bit of a wet blanket, that guy.
One last thing. A sanity check, if you will. I keep plugging the performance numbers in and coming up with a pop-up first stage that can loft 2 times its own GLOW. I see this was already mentioned above in that Vanguard study but that is really amazing.
“Myfundsalow” is an East Indian name, as it rhymes with bungalow
But, I wonder if blaming him makes less sense than blaming an aerospace culture that finds ways to make a tiny fraction of technological solutions so expensive that they gobble up all available funds almost all of the time?
John Bossard:
Dick Morrison was a consultant for Aerojet, and passed away in the late 80’s while in his early 60’s. Tragic. His wife, June, found him sitting in his favorite chair in his den. What was great about Dick was that he loved to talk rockets and other aerospace vehicles, and with anyone who showed any enthusiasm for the subject matter.
Jsuros,
Well, it’s a long and sordid story regarding the use of the ATR for space access applications.
There is actually a rather considerable body of research evaluating the use of the ATR for SSTO, and especially for TSTO applications. The ATR was actually one of the technologies considered for NASP back in the eighties, before the lure of the scramjet pushed it aside. In trade studies done by Boeing even up through the nineties, the ATR offered the lowest gross lift-off weight for a given payload in a TSTO configuration.
I think you make an important distinction when you mention its use for suborbital vehicles. This in fact would be an extremely good fit for the ATR. So, your question about what’s missing is very relevant.
The history of the ATR is really the tale of a new technology trying to complete with and possibly displace incumbent technology. In this case it’s the ATR versus the rocket and the turbojet. This new technology insertion process is well evaluated in Christensen’s “The Innovator’s Dilemma” and ”The Innovator’s Solution”. In competing against their respective strength areas (high T/W for the rocket, high Isp for the TJ), the ATR loses every time (the glass is half empty). It’s only in situations where you need both sufficiently high T/W and sufficiently high Isp that the ATR may be the best solution (the glass is half full). Usually what happens, though, is that when a vehicle designer is faced with such a dilemma, he or she finds a way to reduce the requirements so that a rocket or a turbojet becomes the solution. The ATR loses out once again.
I will share with you my experience in attempting to find applications and users for the ATR. When you approach a vehicle developer about using the ATR as the propulsion system, the high Isp and thrust-to-weight are very attractive. However, when asked when such an engine can be delivered, the enthusiasm gets more muted. When asked about what kind of flight experience exists with the ATR and the answer is zero (no ATR has ever flown to my knowledge), there’s an awkward pause, a thank you, and that’s the last I hear. Vehicle developers want something that’s ready to go, or can be developed very quickly (months, not years).
Let me give you an example. I will tell you exactly what would happen if I approached, say, XCOR, with the ATR concept, and said: “Hey Dan, I’ve been developing a LO2/RP1 ATR engine. Once I get through some more development programs, I should be able to deliver an engine with a T/W of say around 15, and Isp of around 1200 seconds. Would you have any suborbital vehicle applications that could use such an engine?” Dan would probably say something like, “Well, that’s great. But we’ve already got a pretty good LO2/RP1 rocket engine, that’s got a higher T/W, and a sufficiently high Isp to do the missions we want. Also, if we were to try and use your ATR, we’d have to redesign the vehicle with inlets, that’s going to change the aerodynamics, C.G., etc. etc. So I don’t think we can use your ATR for our existing vehicles or currently planned vehicle concepts”. And Dan would be correct. To use the ATR, it has to be integrated into the vehicle from the beginning, and the mission constraints on the vehicle have to be sufficient so that the ATR remains the best solution. Otherwise you just fall back to either using a rocket or a turbojet. Furthermore, folks that are used to the relative simplicity of rocket engines don’t like the ATR’s turbomachinery. It’s complicated, it’s heavy, it’s finicky (all a matter of opinion, IMHO). From a vehicle designer’s perspective, going from rocket to airbreathing propulsion is a huge step change. In fact, its probably too big to swallow. It’s much easier to go from airbreathing to rocket propulsion, which is basically the well-developed path of adding a rocket motor to an aircraft.
In summary, when vehicle developers are forced to choose between the promise of a currently-under-development ATR engine and an existing rocket engine, they choose the rocket. This is why the use of the ATR has languished.
There is also the issue of a general lack of familiarity with the ATR cycle. Perhaps if more folks were aware of the advantages that the ATR offers, there might be greater interest in it. But to truly appreciate the ATR’s advantages requires a little bit more understanding of the design constraints that go into developing a vehicle that can support a particular application. It’s not simply the the highest Isp or the highest T/W, it may be some local optimum somewhere between.
This having all been said, I still believe in the ATR, and I believe that given the right mission and customer, that the ATR will find its place. That’s why I’ve continued to develop prototype ATR hardware, and continue to look for appropriate missions and interested customers. If you think the ATR offers some advantages, then get familiar with the cycle, and help spread the word. As we say in the (extremely) small ATR circles, the ATR is still a solution looking for a problem.
John B hit the nail on the head here. The problem is that an engine “good enough” for suborbital purposes can be designed and tested out by a team of 4 or 5 guys over the course of a few years (with several iterations useful for smaller vehicles along the way). Doing an ATR would be great technically, but since it has turbomachinery, it’s going to cost a bit to develop. Now, as a “strapon booster” or a reusable first stage? It might make sense. But the fact that it’s hard to build a cheap subscale one that you can quickly use on an early suborbital booster does make it harder to adapt it for use.
John, just out of curiosity, if you were trying to do a LOX/hydrocarbon, 2500lbf ATR, how expensive do you think it would be to do one? One nice thing about our vehicles are that while we’re trying to keep the core technologies simple and low-tech, the vehicle is designed from the start to allow it to be used as a testbed for stuff like this.
~Jon
Jon Goff,
Because of my interest and belief in the potential of the ATR, I embarked on my own ATR “development” program a number of years ago. That effort culminated in the hot-fire testing of the world’s first bi-prob GG ATR. The pictures on my blog show some of the results of that effort. The core turbomachinery was from an automotive turbocharger. I used both the compressor and the turbine from this system, but designed, and fabricated a new compressor diffuser, combustor (afterburner), gas generator, and turbine manifold. My friend John Bergmans (Bergmans Mechatronics) designed and built the LABVIEW based instrumentation and control system. John and I worked most weekends for several years to design and test this engine.
Because ATRs do not require the high compressor pressure ratios that turbojets do, they usually only require a single compressor stage, and are more forgiving for off-optimum performance. All of this translates into lower design and fabrication costs for ATRs, compared to turbojets.
Any cost estimates I would provide would be largely guesses without a more careful assessment, but let me give you a few ATR cost numbers that I can honestly vouch for. The ATR development effort indicated above probably cost me somewhere 20k to 40k, including building the test pad (and no labor costs!). This engine is just a test bed engine designed for anchoring prediction codes, and evaluating relative component performance. Predicted performance was about 80 lbf static thrust, at an Isp of around 1200 seconds.
Based on cost estimating work for tactical missile ATRs, I think we figured that an ATR might cost 40% to 60% of the cost of an equivalent (by thrust) turbojet, and be about half the size. For sake of comparison, the GE J85 makes around 2900 lbf static thrust (non-afterburning). A new J85 is probably around $1M. So $400k to $600k might be as good an estimate as any (You can ask Mike Melville at Scaled what they paid for their J85s when they bought them off ebay )
When you start talking turbomachinery, people start getting jittery because of the perceived complexity and cost. It’s certainly going to be more than a pressure-fed rocket, but it can still be manageable. Perhaps we can talk more about a flying ATR test bed. To my knowledge, no ATR has ever been flown, could be ground-breaking.
John B: a question:
My understanding of ATRs is primitive, as I thought it was purely an airbreathing engine. I didn’t realize it had an oxydizer combustion stage as well. I’ve heard the ATR spelled out as an Air Turbo Rocket, and also an Air Turbo Ramjet. Are these designations one and the same, or is the “rocket” version the one that has a LOX staging as well?
I would imagine that the most compelling use for ATRs in a launch vehicle would be as the flyback stage of a TSTO system. It seems a more elegant solution than a rocket-propelled flyback booster. It would seem also to be easier to design as a flyer.
I would imagine that an ATR would be best suited for a parallel-burn TSTO, especially a fairly large one.
Roderick,
You are correct. The ATR is indeed an airbreathing engine. And don’t feel bad about being confused about the exact nature of the cycle, and even it’s name. For Historical reasons, the TLA (Three Letter Acronym) for this particular engine cycle is the ATR. It can mean “Air Turbo Rocket”, as well as “Air Turbo Ramjet”. They are used interchangeably, which is why I generally just stick with ATR. The term ATR also applies to the regenerative (or expander) cycle engine as well.
Ironically, it is the very nature of the engine itself that has given rise to not only the confusion of its name, but also has contributed to its lack of popularity, and perhaps even more fundamentally, to the actual mischaracterization of the engine itself.
The basic reason for this confusion, is that the ATR possesses components from the turbojet (the compressor, the afterburner), and components from the rocket (a fuel-rich gas generator, and an impulse turbine (as opposed to a reaction turbine) ). The turbojet guys didn’t like the ATR because it had rocket parts. The rocket guys didn’t like the ATR because it had turbojet parts. The ATR is the red-headed stepchild, it is loved only by the outsiders in each great propulsion tribe.
Ironically, it was the rocket guys that took up the mantle of the ATR. Engineers at Aerojet (a rocket house) did much of the early development work on ATR cycles. However, the largest ATR development work was done at Pratt and Whittney, West Palm, the fabulous model 304 hydrogen regen ATR, dependable engines indeed.
When rocket guys design rockets, they know that the highest temperature, and thus highest exhaust velocity (and thus highest Isp) occurs near the stoichiometric mixture ratio of the propellants. Thus, it was assumed that the highest Isp for the ATR would occur at near stoichiometric conditions in the combustor. This turns out to be dead wrong. The Isp of the ATR monotonically increases as the combustor is leaned out. Now, the thrust goes down as well, but the flow rate decreases faster than the thrust decreases, thus, the Isp increases. This fallacy was not discovered until the 1990’s by Dr. Kirk Christensen (my esteemed colleague at Aerojet). We call this misnomer “the rocketman’s folly”.
But the turbojet designers were not immune to errors either. In designing turbojets, you generally assume that the all-important gas properties entering the turbine, gamma, Cp, are pretty much close to that of air. For an ATR, this turns out to be dead wrong. The gas properties entering the turbine are whatever is coming out the gas generator (or heat exchanger, if you’re a regen cycle). Ideally what you want is a low molecular weight, and a high gamma. (That’s why hydrogen is so great). We call the misnomer “the turboman’s folly”. Even more ironically, either party, rocketmen or turbomen, would have readily corrected the mistake of the other. But we don’t communicate like we should.
The bottom line, is that the ATR, despite its inherent simplicity, has its own issues.
Regarding launch vehicle applications, an airbreathing first stage to a TSTO system is indeed one of the most compelling applications for the ATR. Trade studies by Dr Jim Martin of Boeing showed the ATR to airbreathing propulsion engine of choice. Bill Escher, the patron of combined-cycle engine work in the U.S. since the 60’s, also found the ATR to trade well for space launch applications.
However, I believe the truly most compelling application for the ATR, is that in which the ATR is integrated with a rocket motor. Such an engine would be capable of producing static thrust, flying through the sensible atmosphere, and then transition to rocket propulsion for exo-atmospheric propulsion. I have termed such a concept a “transition engine” (try googling the term). Such an engine would transition continuously from airbreather to rocket. The very elements that have been problematic for the ATR, i.e. its considerable overlapping of turbojet and rocket parts, are the synergy that makes such an integration between an airbreather and a rocket engine possible, and makes the concept of such a transition engine possible.
ATR engines sounds like part of an aircraft to replace the U-2 and Blackbird spy planes. The unmaned RQ-4 Global Hawk is useful for tactical photographs but its slow speed make it vulnerable against an enemy with anti-aircraft missiles.
John B:
How would the operation of an ATR change after it transitions to all-rocket for exo-atmospheric propulsion?
Roderick:
the operation, and performance, of the rocket-mode of the previously described “transition engine” would simply be that of a rocket motor, possibly with some modest compromise to its performance caused by having to integrate it with another system.
I hesitate to go into much detail on such concepts, because they are highly speculative. However, other transition engine developers, such as Alan Bond, are more sanguinary about their potential performance, so I guess I can be too.
Thanks, John B. for the feedback!
John B:
I assumed that the full-rocket mode of such an ATR would not have the high performance of a traditional launch vehicle rocket engine. Assuming that I am correct, would the full rocket mode still enable orbital velocity? I’m making a second assumption here: namely that the advantages of an air-breathing boost phase might cancel out the lower performance characteristics of the follow-on rocket phase.
I recognize, as you stated, that this is all speculative, but was wondering how much you had thought out the concept to date.
Okay, an IANARS question here -
Can somebody give a brief outline of the differences and proposed relative advantages between this concept and the SABRE idea from Reaction Engines?
a good idea that will NEVER become a “cheap” idea! “save” a few tons of LOX never will be cheap enough to worth the costs to develop and build a “turborocket” and the same will happens with the UK Skylon (that has several other design flaws)
I decided to let Gaetano’s comment in this time (because for once it’s actually on topic, and even though rude, poses a reasonable question). The question is, can an ATR really reduce overall costs compared to an all-rocket system? The fact that’s often pointed out is that costs tend to scale more with dry mass than with GLOW, and while an ATR will likely save you a ton on overall GLOW, you’re mostly replacing cheap tankage and LOX with a much heavier engine….now admittedly it’s also an engine that doesn’t have to run at the kinds of absurd pressures that most booster rockets do these days, but it is a legitimate question. At the end of the day, for most applications costs, services, and reliability matter far more than the technical details of how you get there.
Thoughts?
I’ll take a crack at the question of whether an ATR can actually produce a benefit from the customer’s point of view.
As Jon points out, the customer really only cares about the service that’s provided and the cost. Even if it doesn’t reduce costs, an ATR may be able to win in the service category.
Customers want to fly from locations conveniently located near population centers. Existing airports are an obvious choice for that. But as regulations evolve, rockets flying at airports may be required to have minium loiter time and divert distance capabilities that an ATR can provide, but a rocket cannot.
Speaking of the customer’s point of view, I was wondering if a ATR powered first stage could lift an upper stage at a much more leisurely rate than an rocket, which must get out of the atmosphere as quickly as possible to minimize fuel usage.
If you design a reusable ATR first stage and “Pure” Rocket second stage with equal GLOW the first stage would have a lot of delta vee margin to expend on things like staying subsonic below 60,000 feet and propulsive braking on reentry below 120,000 feet. All benefits from the point of view of passengers and spaceport neighbors.
Jon,
Scaling John’s BOTE cost estimate above up an order of magnitude gives us an ATR of RL-10 level thrust for 5 million plus or minus 1 million. This compares reasonably with a 3.5 million stock RL-10.
So, for twice the cost of engine you can get a first stage with a GLOW equal to rather than twice that of the second stage and a lot of delta vee margin that could help with reusability.
Hm. The maximum engine thrust you need in the first stage scales with total GLOW of the vehicle, doesn’t it? So you end up paying twice as much per unit thrust for a smaller ATR engine needed to lift your second stage out of the atmosphere. Lox and tanks may be cheap but you do have to pay for extra rocket engine to lift them.
“”"Can somebody give a brief outline of the differences and proposed relative advantages between this concept and the SABRE idea from Reaction Engines?”"”"
Erik W:
I can only give you a partial response. The SABRE is a more complex design in that it is collecting additional air or oxygen while still in air-breathing mode to help oxydize the rocket mode. “Son of HOTOL” if you will.
I believe, and John B. can correct me on this, here, when we are talking about the latest iteration of this thread that talks about a rocket mode, the ATR does not attempt to collect additional oxygen, and is a simpler configuration than SABRE. It would rely on already-stored LOX for its rocket mode.
My numbers suggest that the turborocket will find it hard to beat rockets in the verticle launch mode alone. They can do service as launch assist and landing with loiter margin on VTVL RLVs. If the required hover exceeds a minute or so, ATRs can win on hardware mass.
As John B mentions here and on his blog, the ATRs win when one vehicle needs both cruise and acceleration performance. The projected rocketplane vehicle with both jets and rockets could have been well served by a developed turborocket. Either to replace both, or to drop the turbojet mass while extending its’ mach range.
One type flight profile that suggests ATR use is a horizontal landing fly back booster that operates from multiple airports with noise restrictions and limited corridors for higher speed flight. Say one that takes off from a runway and must stay under 250 knots until it is 30 miles from the airport and has over 10,000 feet of altitude. This would be reasonable distance from the population centers that airport serves. This part is flown on a medium-high Isp mode with low propellant consumption compared to rockets, though high compared to turbojets. At the designated point, it accelerates at maximum thrust/minimum Isp with very moderate consumption compared to rockets, very high compared to turbojets unless they are on afterburner. From mach 3 to 6, turbojets become useless, and consumption is still low compared to rockets. At mach 6 and 120,000+ feet, the second stage separates and continues to orbit. The flyback vehicle reenters 300 or so miles down range and must fly home. It can co that by operating vey lean and almost approaching turbojet consumption levels. In this profile, neither the rocket nor turbojet could do the job alone, and using both really ups the mass.
Another potential profile is an air launched fly back stage. Gary suggests that 1,000 mile range is apropriate for air launch to maximize flight opportunities to a given orbit. An ATR vehicle would weigh less than the rocket vehicle and could be more compact, both of which reduce carrier aircraft difficulties and increase potential payload. The stage is released and accelerates to mach 6 over 120,000 and releases the orbital stage. After reentry the vehicle must fly 600-800 miles home .
Another is an airlaunched fly forward verticle landing stage. The carrier aircraft flies to a spot uprange that lets the booster reenter near its’ home base. The ATR could provide serious cross range compared to rockets to correct for missed reentry targets. Also landing margins could be relatively large.
John B wrote:
>The ATR has indeed been proposed as a strap-on type booster.
>This may be a viable configuration. You couldn’t use turbojets in
>this fashion because their thrust-to-weight isn’t quite high
>enough, and ramjets and scramjets don’t make static thrust.
>For this application, the ATR is really the only viable airbreathing
>engine that would work.
Sorry to disagree with the expert, however… The low-bypass turbofan F100 engine actually DOES have the T/W to carry this operation off. Especially in full afterburner mode, but the thrust of the latest model varies from over 17,500lbs dry (non-afterburner) to over 29,000lbs wet or full AB mode. Current T/W is 7.8:1 in most applications. (The F100 was the first production turbofan/jet engine to achieve a positive T/W and gave both the F-15 and F-16 the ability to climb AND accelerate verticaly without the use of afterburners. The newest upgrade to the F-15 will replace the F100s with twin F110 engines.)
The engine used in the F-14, (and soon the F-15) the F110 has even better performance, 27-28,000lbs dry, and over 30,000lbs in full AB. Lifetime improvements for the next block of F110 engines is expected to base-line at around 30-32,000lbs with proposed modifications. The current engine T/W is 6.36:1 with upgrades it’s expected to achieve over 7:1 at the very least.
Continued;
>It is the ATR’s intrinsic simplicity that allows it be much like a “can”
>strap on.
The strap-on “can” idea came from the original work by Edar IIRC, a lot of work done by Glenn Olson, (originally available at a dedicated website alt-accel.com, but since had to be taken down due to financial issues, see: http://www.pulse-jets.com/phpbb3/viewtopic.php?f=4&t=5201 but some work on the Tri-Mode Amateur Rocket Launch Assist, and the POGO jet-powered first stage concept can still be found at: http://home.earthlink.net/~altaccel/ ) on jet and ramjet assist showed that the strap-on approach is self limiting due to vehicle aerodynamic interactions and that a more probable and likely ‘better’ method would be a booster stage application. Edar’s “booster” jets cut off at 50,000ft and around Mach-2 for several reasons, those stated included limiting the full-afterburner run time to 60 seconds, and being able to recover the purely ballistic engine “pods” within 10-20 miles of the launch site on dry land.
(Mentioned at one point was the launch point being situated in “East Texas” to take advantage of large “ranch” lands as a launch and recovery site for the jet engines. The two-barrels of the “first” rocket stage would re-enter and land off the tip of Florida while the third-barrel “second” rocket stage would continue to orbit and then re-enter and recover at the launch site)
Not mentioned by Edar, but found to be a major factor in using air-breathing boosters or a booster stage by Olson was that after Mach-2 standard turbo or ramjets began to require careful “shaping” of the upper payload to continue to feed air into the intakes, in addition to variable intakes and ramps to continue to drop the intake air to subsonic speeds at the compressor face.
As I understand it, (please correct me if I’m wrong) the ATR would probably not be as effected by these issues?
>The April 1959 issue of Astronautics has an interesting article on
>using the ATR as a first stage booster for a Vanguard-sized small
>satellite launcher. Considerable reductions in overall gross lift off
>weight were realized by using an ATR instead of a liquid rocket
>(22,000 lbm for Vanguard vs. 11,300 lbm for the ATR-powered first
>stage).
That’s quite interesting as studies have consistantly found that reducing gross take/lift-off weight of the launch vehicle directly correlates to cost reductions in manufacturing and maintenance, as well as development time and cost.
>And yes, our studies showed that an ATR could reach altitudes in
>excess of 120,000 ft at supersonic flight speeds. At a given flight
>speed, the ATR can (theoretically) fly at higher altitudes than any
>other airbreathing engine. This is because it retains the highest
>pressure in its combustor (for a given compressor ratio, or inlet
>compression ratio for ram/scramjets) compared to any other
>airbreather (see plasma wind blog post, part 1)
Nice. But the question arises again, is such performance worth the extra cost in vehicle to protect from heating issues from protracted high supersonic/hypersonic travel even at those altitudes?
Having asked that I’ll note here that being able to “loiter” at such altitudes even as subsonic speeds would be a great boon for commercial operations of a Horizontal Takeoff/Landing vehicle as they would remain outside “normal” air traffic patterns disconnecting themselves from strict takeoff/landing times needed for Air Traffic Control as compared to orbital entry and reentry windows.
Hopefully I can resolve my issues with your blogs “comments” section as I’m sure everyone will shortly be tired of my comments and questions here
(Better to inflict them on you on your home turf as it were)
Randy
“”"”"Edar’s “booster” jets cut off at 50,000ft and around Mach-2 for several reasons, those stated included limiting the full-afterburner run time to 60 seconds, and being able to recover the purely ballistic engine “pods” within 10-20 miles of the launch site on dry land.”"”"”"
I like this item, Randy, because I’m a big fan of short-duration boost phases, assuming you have a high-performance orbiter stage. I always found it perplexing that SSTO designs weren’t scaled back a bit when they ran into the inevitable mass-fraction barrier. An SSTO with unacceptable weight penalties becomes viable with a modest boost that allows easy recovery of the booster(s), so it’s still operationally simple compared to conventional llaunchers.
I’m sure everyone here is familiar with the all-ATR pop-up booster concepts, where the primary vehicle is the ATR-powered one, and a self-propelled payload is released, OR, the payload is lofted to intercept a skyhook tether dangling from an orbiting station.
Would Mach 6 and 120,000 ft. be sufficient for such a concept? I envision “pop-up” as meaning a vehicle attains a very large fraction of orbital velocity and altitude (a lot more than Mach 6 and 120K ft.), so maybe it only makes sense as a “tether interceptor.” Thoughts?
And, is it realistic to push ATR speeds beyond Mach 6?
Roderick,
My understanding of a “pop-up” booster comes from the book “The Rocket Company”. In this idea the first stage is used to lift the second stage out of the atmosphere and give it all the vertical velocity it will need to avoid gravity losses as it accelerates to orbital velocity. The second stage gets no help from the pop-up stage in reaching orbital velocity, but it does all of its acceleration in vacuum and wastes no effort fighting gravity. Think of it as an assisted SSTO concept.
In the book, the first stage lifted the whole vehicle to 100,000 feet and 3000 feet per second (less than mach 3), then fell up out of the remaining atmosphere while firing “sustainer motors” to add 3 m/s/s for another 44 seconds before staging at 200,000 feet. The second stage then had a period of two minues or so in which it could ignore gravity while accelerating up to orbital speed in vacuum.
As I understand the discussion of the last few days, an ATR powered first stage could do everything a pop-up booster requires and help a little with gaining orbital velocity also.
Wow, there’s some really excellent discussion, here, the best I’ve heard in years regarding ATRs. I apologize in advance for the long response, but y’all are worth it.
Roderick:
As much as I am intrigued by the concept of an integrated ATR and Rocket (known an as ATRR back in the 1980’s), it is a highly speculative (and controversial!) notion. So I’m going to have to defer this question, a least for a while. But the point you make is important: does operating as an airbreather provide a net benefit to a orbital or suborbital system against the compromises that it engenders? I think it can, but there are many assumptions and disclaimers that need to be acknowledged.
And your last question: ATRs are going to be hard pressed to get to Mach 6. I see that as really the upper flight speed of this engine. But, maybe it doesn’t really make sense to try and fly faster than Mach 5 or 6 with an airbreather. (see further comments below).
Eric W:
The short answer is that the ATR is a very simple engine, whereas the SABRE engine is quite complicated, but that’s because the SABRE functions in both airbreather and rocket modes.
The SABRE engine, as created by Alan Bond, as gone through many iterations. The very best description that I saw of the predecessor engine to the SABRE, which was the RB545, was presented in Spaceflight Magazine, Vol. 35, May 1993, pp. 168-172, in an article entitled “HOTOL’s Secret Engines Revealed”. Here, the entire engine cycle and schematics are presented and discussed. Now, this was the engine concept from 1993. I believe that considerable simplifications have occurred in ensuing years, I and believe that Bond has publicly discussed some of these developments. You can see some of the resulting simplifications by looking at a schematic of the SABRE engine which can be found here:
http://www.reactionengines.co.uk/downloads/JBIS_v57_22-32.pdf
Jon Goff:
The question of realizable cost reductions for any particular vehicle configuration is, simply put, very difficult to answer definitively (I really think you may know quite a bit more about this than I do!). So, I don’t think I can provide an answer that will be any better than anyone else’s. It’s good that you pointed out that costs tend to scale with dry mass. They also scale quite well with parts count!
The common refrain is that airbreathers don’t buy you anything because LO2 is so cheap. Just make the vehicle as big as you need, and use as much propellant as you need, because propellant costs are a negligible recurring cost for launch vehicles. But there are some subtleties that get neglected in these considerations.
First, the overall development cost IS strongly influenced by vehicle size. Big vehicles require big facilities, and big everything. One of the rules that Burt Rutan lived by in making SS1 was to keep it as small as possible, and that’s one of the reasons for his success. The second point I would offer is one that we have not truly encountered to date. In mature transportation systems, propellant costs asymptotically approach about 1/4 to a 1/3 of the overall operating costs (it is said, I have no reference to offer here). Airlines go to a great deal of trouble to find ways to reduce fuel consumption. A reduction of a less than a percent in fuel consumption over the fleet can translate to millions of dollars in savings for an airline. Someday, when launch vehicles blacken the sky, we will truly care about propellant costs, and then the concomitant savings in propellants because you have a somewhat smaller vehicle, and you have somewhat better Isp, will matter. Perhaps it may be the very advantage that makes your space freighter company more competitive that someone else’s who is without those advantages.
But, that’s a long way off, and so perhaps it’s premature to ask that question. Jon, a while back you wrote about what’s the appropriate technology to be working on right now. I thought it was a very balanced piece. Maybe right now, airbreather’s don’t make the best practical sense for current or even next gen launch vehicle concepts. But if we’re collectively successful, then maybe they’ll be a time when we’ll want, nay, demand airbreathing systems to provide reduced propellant fractions. Somebody’s gotta’ be working on those systems today so that they’ll have a better chance of being ready tomorrow. I’m one of those guys: lonely, unpopular, with nothing but a well-converged solution to keep me company. But someday I believe, airbreathers will find their way into the cohort of viable launch vehicle propulsion options.
Bob Steinke:
I’m glad you brought up the point about serviceability and practically. I think it’s another important, but underappreciated aspect of suborbital and orbital systems.
Jursos:
Airbreathing launch trajectories are indeed different from that of rocket-powered launch trajectories. In fact, finding minimum propellant-usage trajectories for airbreathing vehicles is a relatively-underexplored area, and it has the potential for significant improvements. Dr. Frank Chavez at AFRL has done and is doing some great work in this area.
John Hare:
ATRs have definitely been identified as being a good candidate for fly-back boosters. And as previously discussed, as an airbreathing propulsion for the first stage for TSTO concepts.
I’ll have to give some thought to the fly-forward concept you mention, interesting.
Randy Campbell:
Vertical launch vehicle concepts are somewhat sensitive to booster T/W. I don’t have a hard number for what kind of T/W you’d need to make a concept viable. My guess is that you’d need a T/W that’s 10:1 or better, but I could be wrong.
Thanks for the references on strap-on’s, I’ll check those out.
Inlet are a big issue for any airbreather. And while the ATR is indeed less-susceptible to inlet distortion and unstart issues, it’s performance can still be deleteriously affected by these issues.
The NASP work done in the 80’s showed that the advantages of a scramjet engine were more than offset by the additional thermal protection one would need to plow through the air at Mach 10-20+. I think what we learned is that there really is a practical upper bound on the flight speed of airbreathing engines that makes sense. I’ve seen this upper bound gradually creeping down over the years. My hope (and suspicion) is that eventually, that upper bound might end up at about the upper operating limit of the ATR! (about Mach 5-6).
To my knowledge, no one has ever mentioned the advantages of being able to loiter at high altitude in terms of air traffic control, so that’s a great idea. Keep in mind that an ATR would need to be flying fast (Mach 2-3) at high altitude to stay within it’s operating conditions, but it’s still a practical advantage.
HI John. I noticed that John Bussard mentioned on his website that one of the advantages of the ATR over the plain turbojet is that the power to the compressor is independent of the airflow. Since your design has some of the airflow passing through the gas generator, isn’t this advantage negated? Or at least reduced, given that the cooling provided by the LOX injection allows a higher speed for a given turbine inlet temp(?). I wonder if a dual-mode engine that can divert all the airflow away from the combuster to allow a higher speed would be practical.
Reading over John Bossard’s comments about the desirability of a transition engine. I think I have the beginning of an idea for such an engine, but it would take too long to describe, and also I don’t know how to post pictures. Could I email you about it?
John B,
In John H’s previous posts involving rockets with turbines, I’ve been voicing a concern about having moving parts inside the thrust chamber as “not such a good idea” due to combustion instabilities and so on. Obviously you’ve been able to get something like this to work. Are my concerns valid? I guess what I’m trying to ask is: Are their considerations for conditioning the combustion products (perhaps redirecting through some vanes) prior to them being used to drive the turbine? I have a CFD background, but no turbo-machinery specific experience, so I’ll defer to you on this point.
Tim.
spamharedotjohnatrocketmaildotcomspam
It has become my opinion that an engine with some of the charictoristics of the ATR and turbojet would be better for the purpose than either alone. It will take me a while to do a proper post on it with a few numbers. I have a hint of an idea that some ATR concepts could be incorporated into commercial turbofans to reduce kerosene consumption. Sometimes the answers come quickly, and sometimes not at all. If a general commercial or military engine reached production, bolting one on for space use would become an easier decision.
Roderick Reilly wrote:
>I like this item, Randy, (Edar Jet-Booster concept) because I’m a
>big fan of short-duration boost phases, assuming you have a
>high-performance orbiter stage. I always found it perplexing
>that SSTO designs weren’t scaled back a bit when they ran into
>the inevitable mass-fraction barrier. An SSTO with unacceptable
>weight penalties becomes viable with a modest boost that allows
>easy recovery of the booster(s), so it’s still operationally simple
>compared to conventional llaunchers.
Thanks Roderick, I keep running across rather ‘obscure’ concepts and am always happy to pass them on and keep them going. Annoying, but a personal habit I can’t seem to break
)
My experiance over the years has taught me that SSTO advocates tend to feel that it’s “Single-or-Nothing!” in that they will REQUIRE the launcher to be a SINGLE stage or not support the concept at all. Given that they will then usually resort to all sorts of verbal-gymnastics to down-play any “booster” concepts that might help make the vehicle actually work. The will talk of having “zero-stages,” (which btw is actually funny as the concept of the zero-stage is technically so that an SSTO vehicle arrives in orbit with more on-board fuel left for other mission needs but is now pretty much the ‘cheat’ name for a SSTO stage booster
) “assist-rockets,” “launch assist vehicle/platform/etc” and other ‘names’ have been applied to get around the idea that a “SINGLE-stage-to-orbit” vehicle might not actually be JUST a single stage. 
(And just an FYI Jon, I’ve seen the concept of an air-launched SSTO “re-named” as an SSTO with airborne mobile launch platform
John B.
The Edar concept used 10 mil-surplus high-time F100 engines with combined lift-off thrust of 290,000lbs under full-AB. In addition you had 6 under-expanded RL-10 engines, which would come out to somewhere under 200,000lbs thrust for lifting a launch vehicle with a total GLOW of about 145,000lbs. I found a link to the concept here:
http://yarchive.net/space/launchers/jet_first_stage.html
Mach-6 is plenty fine with me for the ATR, that’s great for the sub-orbital Type 1 and 2 missions I’m thinking of.
)
(BTW: Just an FYI John, and a reminder to everyone else: In order to help with discussions on a couple of email lists and other discussion forums I’m on “we” came up with a series of ’short-cut’ desctiptors for various flight profiles which I tend to use everywhere now. In reference to the above there are two distinct “sub-orbital” flight missions described as “1″ and “2″ as above. Sub-Orbital Class-1 missions are “X-Prize” flights involving little more than vertical flights to “space” and back with little or no horizontal distance covered. Sub-Orbital Class-2 missions are more “point-to-point” sub-orbital in that there IS a horizontal component that can be quite large. Class-3 is orbital missions and we generalize Class-4 as Lunar and Beyond at this point. We now return you to your regular comments section
Getting out of the atmosphere after Mach-6 would keep your materials costs down and well as your structural needs. My thought would be to combine the ATR with a body shape capable of some external burning of fuel and using a ‘duck-and-drake’ skip-glide trajectory. The external burning of the fuel has been shown to extend the skip-glide with minimum energy and to aleviate the high heat loading and structural loads during the ’skip’ portion of the flight.
On an ATR with a more effecient “rocket” portion; I’ve seen patents and information on several cycles of proposed air breathing “rocket engines” around the net. As an example, one concept used an RL-10 engine for the rocket with the cold hydrogen being used to turn a turbine as it expended and the LH2 and compressed air being fed into the combustion chamber of the RL-10 until such time as the RL-10 was switched to pure rocket mode. Since it didn’t have an ‘after-burner’ or direct attachment ducting into it the difference seems to come down to how the ‘rocket’ part fits into the description. I’d like to hear your comments on the differences?
As to flight testing an ATR, you said that your current ATR work used automobile turbocharger components which would make it fairly ‘cheap’ for a small team to do. Have you ever heard of the “Forerunner” concept home-built super/hypersonic aircraft?
)
(I’m guessing probably not as most folks I ask haven’t
While not really more than a ‘thought-experiment’ really the overall idea can be found here:
http://www.dcr.net/~stickmak/JOHT/joht04forerunner.htm
(The “Joys of High Tech” section is quite an interesting read in and of itself)
My ‘idea’ here though is something more along the lines of the “EZ-Rocket” in taking an exsisting airframe and substituting an ATR engine in place of the standard one. Actually in the case of the Long-EZ this seems to be more a ‘feature’ of the design since beside putting a rocket engine in one the airframe has been used to air test a working Pulse-Detonation engine, a Prop-jet engine, a turbojet engine (including one made from a truck turbocharger) and a home-made fan-jet engine!
I personally know someone who would probably be QUITE willing to fly such a test aircraft, he and I were originally discussing building and selling “home-built” rocket plane kits so I’m sure he’d LOVE to fly an ATR test bed if you need a volunteer
Randy
Randy:
Thanks for the feedback. I’m a big believer in what I call “Virtual SSTOs,” such as we’ve both posted about here. Another option for a visrtual SSTO is a drop tank, in my opinion. The boost-phase propellant can be contained in a drop tank which can be parachuted into the ocean for retrieval. This idea is nothing new, as a couple of the earliest space shuttle cocepts were just that.
Eric Collins:
The primary issue for turbines is the elevated-temperature strength of the turbine material. The turbine blades in turbojets use active cooling to keep the blades from exceeding their material temp limitations. Rocket turbopumps use fuel-rich combustion mixtures to stay below material temperature limitations, along with other approaches, such as thermal barrier coatings. So, I would say “yes” to your question. You run your gas-generator at mixture ratios that allow the products of combustion to be at or below your turbine material temperature limitations. I don’t combustion stability as being an issue for rotating machinery in combustion chambers (if I understood your question correctly).
For the ATR cycle, this is fortuitous, because the fuel-rich turbine drive gas is subsequently burned with air in the ATR’s combustor.
Hope that answers the mail.
Randy C:
I was not aware of the trajectory shorthand, so thanks for pointing that out.
I had not heard of the forerunner concept, I’ll check the link.
No one has flown an ATR yet. The person who does that will be the first.
I agree with your comments regarding SSTO. It seems like we got hung up on trying to make this work, and have left little room to consider other, perhaps suborbital or TSTO approaches that might allow us to more incrementally develop aerospace vehicles. I think we should be trying all sorts of ideas, and then let the market decide what’s the best approach.
Just to ’stir-things-up-again’ since everyone else is moving on, (yes I DO have issues with moving on
)
What kind of airframe would (anyone) you think would make a good tech-demonstrator testbed for an ATR?
XCOR used the Long-EZ for the EZ-Rocket, which would probably work for an ATR but I’m thinking maybe something more… Lifting body-ish maybe?
Hmmm, maybe a variant ‘ultra-light’ FMX-4 Facetmobile?
http://www.facetmobile.com/
Randy
Randy:
I like the facetmobile. I think lifting bodies are great because of their more efficient use of structural weight.
For a testbed of a flight article for a first stage ATR vehicle, I think a different design would be needed, since it would have to serve as a carrier for an orbiter stage. Perhaps an aerodynamic “cradle” design that is all engines and aerodynamic surfaces, and virtually no fuselage would be the best bet.
I for one would like to see if SpaceShipThree could make use of something like the ATR. IIRC, SpaceShipThree would be one that they would consider doing point-to-point suborbital hops with. Such a vehicle should also be able to launch a fairly substantial orbital payload (not the vehicle itself, but via a small third-stage rocket carried on board).
I think the toroid tank structure for a lenticular lifting body would be good. The triangle ship I suggested a few months back would be my second choice. Gary Hudson suggests that using the tanks for structure is not worth the trouble though.
Roderick:
We’d not be looking at a test-bed of a first stage, rather like the EZ-Rocket the idea would be to flight test the ATR engines in an aircraft and gather flight data. The Facetmobile “shape” elongated and made more super/hypersonic acceptable would be a choice for “down-the-road” testing at this point I think easy and inexpensive to build would be the more direct criteria.
Though I’ll note here that the FMX has a long Mach “shadow” behind the upper fuselage high-point and along that rear triangle that would suit sheilding an upper stage.
John H:
I was thinking about lenticular also, but most of them seem to do better at higher speeds rather than lower so for an initial flight item I’m thinking a lifting body shape would be better. Speaking of the toroid tanks though, somewhere on one of these hard drives I’ve got a copy of a Boeing patent for a toroid tank concept for a Blended-Wing-Body aircraft. Due to the size of the tank the patent actually shows that the tank structure IS part of the overall aircraft structure so the idea has merit.
Randy
I was interested to read of the ATR concept. However, it still uses the heavy compressors/turbines of turbojets and it also needs oxidizer in addition to the burnable fuel to run.
On another forum I was thinking of ways to get the high compression required for high thrust for jets without using compressors/turbines and not using oxidizer other than that in the surrounding air. I copied this below.
Bob Clark
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The reluctance to use airbreathing engines for part of the time to reach orbit is due in large part to the fact that jet engines are heavy compared to the thrust they can produce. See the list of thrust-to-weight ratios for some engines here:
Thrust-to-weight ratio.
http://en.wikipedia.org/wiki/Thrust-to-weight_ratio#Engines
The thrust-to-weight ratio for turbojets might be only 5 to 6, where as for rocket engines such as the space shuttle main engines might be 73 or above. A big part of this poor thrust-to-weight ratio for jets is the complexity and weight of the compressors and turbines jet engines have to carry:
Jet engine.
http://en.wikipedia.org/wiki/Jet_engine
However, the thrust-to-weight ratio for ramjets because of their simplicity can be quite high:
Ramjet Performance Primer.
“There are no physical limits to the minimum weight of a ramjet other than design and materials. The 1950’s Marquardt RJ43-MA-7 had a thrust/weight (T/W) ratio of about 40. With today’s engineering and materials that could probably be brought up to 150-200 without too much effort. Such T/W ratios would make ramjet powered vehicles excellent accelerators.”
http://www.alt-accel.com/ramjet2.htm
Airbreathing engines need compression of the air to create high thrust. Turbojets use compressors. Ramjets are able to get high compression from the high velocity of the incoming air alone, dispensing with the compressors and accompanying turbines. Then the suggestion is to replace the compressors/turbines in turbojets with other means to achieve this high compression. One method that has been tested is the ejector ramjet, where rocket exhaust is used to accelerate air into the intake of a ramjet, thus allowing the ramjet to operate even at zero speed.
This method still needs to use onboard oxidizer, for the rocket, to operate. An ideal method would only use the burnable fuel to operate, as do ramjets and turbojets. For a turbojet/ramjet intended as the first phase of a SSTO vehicle that uses rockets at the end stage, what might work is to use the very high pressure turbopumps that high performance rocket engines such as the space shuttle main engines use. Since these high pressure turbopumps are needed to be carried along for the rocket phase anyway perhaps they can be used as well during the airbreathing portion of the trip.
There are several ways this might be accomplished. The shuttle liquid hydrogen turbopumps can produce 500 bars of pressure of the liquid hydrogen with a through put of 73 kg/sec each. I’m imagining this high pressure liquid hydrogen be directed into the intake of the jet engine. You want to do this in a way to compress the air. One way might be that the turbopump outlet into the jet engine be in the form of an annular (ring) opening all around the inlet, some distance into the inlet. This would tend to compress the air together as the liquid comes out directed inward to the center. You also want the air to be forced back to the rear of the engine so the liquid hydrogen would need to be angled somewhat also backwards towards the rear.
The liquid would tend to spread out however, and for a large intake say a meter across or more for the large supersonic turbojet inlets, it’s not certain how far the liquid would go to penetrate into the middle portion of the air to achieve the high compression needed here as well, not just the outer air. You want most of the air to be compressed at least to the 20 bar range commonly seen with turbojets in order to achieve the high thrust achieved by the means of compressors.
Then another possibility would be to use an analogue of the ejector ramjet compression method. This works by using supersonic exhaust from a rocket to force the air into the intake, thus being compressed as is the case with ramjets flying at supersonic speeds. Then what we could do with the turbopump’s output, is to use the Bernoulli principle to convert the very high pressure into a supersonic velocity:
Bernoulli’s principle.
Incompressible flow equation.
http://en.wikipedia.org/wiki/Bernoulli%27s_principle#Incompressible_flow_equation
For a streamline at constant height, (1/2)(velocity)^2 + pressure/density = constant. With pipelines leading out of the liquid hydrogen turbopumps of about 30 cm wide, a density of liquid hydrogen of 72 kg/m^3, and mass flow rate of 73 kg/sec, I calculate the flow speed as 33 m/s. Then if we want to convert the pressure of 500 bar = 50,000,000 pascals to high velocity we would get a speed of 1180 m/s, about Mach 3. Then this supersonic flow could be directed into the intakes to accelerate and thereby compress the air as is down with ejector ramjets.
Still another possibility to get the air to flow at high speed to induce similar compression as with a ramjet might be to ionize the air and accelerate it by electromagnetic fields. The turbopumps use a turbine which is a key means by which electric power is generated. They operate at 70,000 horsepower while weighing only about 700 pounds. There is pretty high efficiency conversion of turbine mechanical power to electrical power. However, we would need a lightweight means of ionizing and electromagnetically accelerating the air. A couple of possibilities for the ionization might be by using a microwave generator or electrically charged wires running throughout the inner volume of the intakes.
In any case some of the exhaust from the jet would have to be bled off to run the turbopump. This might seem to reduce the performance of the jet engine but actually quite a large proportion of the power generated in usual jet engines is used just to run the compressors:
What is a Gas Turbine Engine and How Does it Work?
“The cycle that governs the operation of a gas turbine engine is referred to as the Brayton constant pressure cycle. The engine compressor typically requires about 2/3 (!) of the usable heat energy produced in the burner to turn at maximum speed; the remaining energy can then be used to produce thrust or mechanical power, or a combination of the two.”
http://www.turbokart.com/gasturbine.htm
To get an idea of the power we need, we’ll use as a model the J58 engine which powered the SR-71 to Mach 3+. I haven’t seen any numbers on the horsepower generated by the J58 but I’ll estimated it from the 1 horsepower per 2.5 pounds thrust common for turbojets:
Turbojet.
Thrust to power ratio.
http://en.wikipedia.org/wiki/Turbojet#Thrust_to_power_ratio
The J58 generated about 25,000 lbs thrust in usual turbojet mode, so a horsepower of 10,000 hp. Note though that fuel needed to run the J58 is much less than the 73 kg/sec liquid hydrogen put out by the shuttle turbopump. This page gives its fuel use in the usual turbojet mode as 0.9 lb/(lbf-h), i.e., .9 lbs/hour for each pound of thrust:
Pratt & Whitney J58.
Specification of J58-P4.
http://en.wikipedia.org/wiki/J58#Specification_of_J58-P4
This is 22,500 lbs/hr of fuel at 25,000 lbs thrust, or 6.25 lbs/sec, 2.8 kg/sec. This is in jet fuel. Hydrogen would give higher thrust and indeed will use about half the fuel for the same thrust as shown in the attached diagram of turbojet/ramjet/scramjet Isp’s. So this would be 1.4 kg/sec of hydrogen. This is 1/52nd the usual mass flow rate of the turbopump of 73 kg/sec. The power used by a turbopump is proportional to the mass flow rate, so the power needed would be 70,000 hp/52 = 1,346 hp
This about 1/7th the power output of the J58 engine. A problem though is whether this would supply sufficient compression for the high air inflow of the jet. We might need to flow more fuel through the turbopumps than is burned by the engines. But this would mean we are running the jet engine fuel rich. However, the Isp for jet engines is so high we could afford to run fuel rich and still have a significantly better Isp than rockets.
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Bob,
you pose some very interesting suggestions, and I appreciate the time and effort it took to write your comment. It’s especially helpful when someone takes the time to provide references.
I believe your comment deserves a good reply comment in return, and as such, it will me take a bit more time than usual to prepare. But I will try to have a response to you by tomorrow (7/22) if possible.
In the meantime, if you’d like to learn more about some of the details ATR operation and characteristics, I suggest you read the ATR-related blog posts at Plasma Wind. The Introduction describes the general operation of ATR cycles, and the performance attributes and applications posts go into more detail on these characteristics, and how they can be used.
Bob,
You present some interesting ideas. I will comment on them, based on my knowledge base, and perhaps my own opinion. But I would like to preface these comments by mentioning that my comments, or any other person’s for that matter, are not necessarily the final word on the matter. Many, perhaps most, of the great technical achievements of history were made almost precisely because they ran counter to prevailing wisdom of the time. So, listen to some other opinions and use them to help you, but don’t let them stop you from further developing our ideas and concepts if you think you’re on to something.
First I would say, the notion of “heavy” turbomachinery has to be understood in the context of its application. Turbopump-fed rocket engines, for example, may have a lower T/W than an equivalent thrust pressure-fed rocket motor, but their higher Isp more than makes up for their lower T/W. I think the same is true regarding the use of airbreathing engines as an propulsion element in a suborbital or orbital launch vehicle.
Thrust/Weight ratio is quite important in sizing a vertical takeoff launch vehicle. Obviously, you need a T/W greater than 1 (where W is the gross liftoff weight). Typical launch vehicles have liftoff T/W around 2-4. Lower than around 2 and there can be stability problems, higher than 3 or 4 means that you can sustain very high g-forces near you first stage burnout.
Its not simply low T/W performance that is the issue, but the T/W in association with delivered specific impulse (Isp). As an example, consider a “magic” propulsion box that uses no expended propellant, but is able to produce a T/W of, say, 1.1, so it can lift itself, and just a bit more. Although the T/W is low, the Isp is infinite. Such a propulsion system could take you orbit, although the vast percentage of our on-orbit mass would be propulsion system mass. Is that bad?
Regarding the complexity and weight of the jet engines vs. rocket engines, the SSME is quite a complicated engine, and also possesses turbines, and pumps. Perhaps more significantly however, is the fact that, to get its high performance, it operates with very small performance margins, in terms of stresses, temperature limits, etc. That means that the engine is terribly unforgiving for any off-nominal conditions. Small defects in fabrication or assembly can result it the catastrophic failure of the engine. This is why it must be taken off its shuttle, inspected, and rebuild after most flights. This is one of the reasons why its expensive to use. I’ll be getting back to the advantages of margin later.
In terms of actual T/W numbers, the SSME masses about 7004 lbm. At launch, the engine makes about 375,000 lbf (at 100% power-level), giving it a launch T/W of about 53.
Turbojet turbines and rocket turbopump turbines have comparable power densities. It’s really the mass of the compressor versus the mass of the pump/impeller where a good portion of the mass differences arise. Compressors have to do work on gases, which are highly compressible, and take a great deal more energy per unit mass to raise their (total) pressure, whereas liquids are relatively uncompressible , and the specific work to increase their pressure is much lower. Air is bulky, so moving it around and using it to make thrust makes the engine bulky. If you look at turbojets, they take up a lot of volume, and that’s because they are massive air-handling machines. But bulky does not necessarily mean heavy.
Ramjets:
As one of my colleagues used to say “The problem with ramjets is that when they’re sitting at the end of runway, any fuel you put in them just runs out the end”. This is just a throw-away line, but the point is that a fundamental problem with ramjets is that they don’t make static thrust. Contrary to common opinion, ramjets can in fact make thrust at subsonic flight speeds, sometimes as low as Mach 0.7 (no reference), however they have terrible specific impulses, subsonically, and they need to get up to around mach 2 or 3 before they begin to operate efficiently. This inability to make static thrust is a fundamental liability of ramjets. They always require some type of booster to get up to speed, and this represents a significant logistical impediment to their use. Thus, some other propulsion system has to take the ramjet from zero, through mach 1, and up to mach 2 or 3 before they begin to function. Ironically, rockets are usually used to boost ramjets (and scramjet, too). Rockets have their worst performance making thrust at sea-level/static conditions.
There are some other, more subtle issues with ramjets, as well. Ramjets rely on their supersonic inlet to convert flight speed into total pressure. Air that is shock-compressed in the ramjet’s inlet is then fed directly into the ramjet’s combustor. This has the consequence that ramjets are very sensitive to any disturbances to the incoming airstream. Changes in angle of attack, yaw, and/or roll, as well as acceleration, and even atmospheric conditions, can all deleteriously effect the incoming airflow. These disturbances can easily be sufficient to cause an inlet unstart, in which the shock structure on the inlet is disgorged, resulting in a flameout in the combustor. Even without an unstart, flow disturbances can induce flameouts. And while the basic elements of a ramjet are simple and lightweight, the fuel contoller for a ramjet is anything but. Fuel input into the combustor is a careful balancing act between the resultant combustor pressure and the incoming air stream. Too much fuel, you get a flameout, too little fuel, you get a flameout. The net effect is that ramjets are quite finicky, and require a careful start-up process, and cannot accommodate rapid changes in flight conditions. Is this a limitation for launch vehicles? Maybe your flight trajectory can be designed to be “mellow”, but you certainly will have to put up with much-reduced performance margins.
For sake of comparison, I think its worth asking the question: Why aren’t ramjets in use today?”. We’ve had a number of ramjet-powered missile systems: BOMARC, terrier, Taos, to name a few. The basic answer is that ramjets required a boost function, which is usually accomplished with some sort of rocket motor. Since you needed rocket propulsion anyway, and rocket powered missiles could launch themselves, they were thus logistically easier and less expensive to deal with, even though they have significantly lower Isps, and had to be bigger to make a given range. Solid rocket motors are also a lot more forgiving in terms of logistics, and this gets back to performance margins.
I must disagree with the statement that the T/W of ramjets “could probably be brought up to 150-200 without too much effort”. First, I doubt that these T/W’s are achievable (no reference), and second, anytime you’re dealing with airbreathers, its always a lot of effort. The testing alone is a huge effort: supersonic wind tunnels, supersonic flight tests, etc. I don’t know if you ever saw Marquart’s facilities, which were located right at the Van Nuys Airport in LA, but they were very impressive, in terms of the compressors, steam ejectors, air vitiaters, and the general air flow handling equipment.
Ejector ramjets, in which rockets are embedded in the combustor section of the ramjet, are in fact, a viable approach, and a number of different programs have been and are investigating these concepts over the years. Aerojet, for example, has been working on and off for nearly two decades on such an ejector rocket concept, which they called the “strutjet”. MSFC also supported this related work under a program called “ISTAR”, I believe. Note however, that it is the rockets that are providing the static thrust, not the ramjet. An element of the ejector ramjet is that the rockets can also induce or “eject” an airflow into the ramjet, thus effectively improving the thrust output of the system, and increasing its Isp. References that I have seen indicate that with a good ejector design, additional airflows of around 10% of the rocket flowrate should be possible. So this represents a usable improvement.
I believe that you are also suggesting that an injected spray of liquid (a liquid fuel and/or an oxidizer) may also be used as a compression mechanism for airflows. Anyone who has turned on the shower and felt the breeze created by the shower spray knows that the movement of sprays can certainly induce airflows. (My doctoral work was focused on liquid and fuel sprays, so I like sprays).
The possibility of pumping or compressing air using sprays gets down to the basic process of how one can actually pump or compress a gas. A compressor compresses a gas by using pressure body forces to mechanical push the air molecules together. A spray process “compresses” a gas, or more specifically raises the total pressure, by using shear forces between a moving droplet of liquid and the adjacent air to accelerate this air. Once the air has some speed, you can convert its speed into (total) pressure by allowing it to deaccelerate, or diffuse. Using shear forces to accelerate a gas is not particularly efficient. The greater the velocity difference, the more energy is lost or wasted, so the compression process becomes asymptotic. The net effect of all this is that spray “pumping” can never approach the compression ratios that you can achieve by using pressure body forces, such as those exploited in compressors.
In my opinion (and I could be wrong), it will not be possible to use high pressure liquid injection to “shear-pump” air to high pressures. In fact, if you inject the spray into the inlet with a high enough velocity (supersonic droplet velocities, for example), you start to become less of a “spray-compressor” and more of a liquid-jet cutting tool. The droplets can easily have enough energy to knock a hole through your inlet walls. I have never seen any data that suggests that you can achieve 20 bar pressure rise using spray pumping. But I could be wrong, and maybe there’s some data out there on it. If you find some, I would sincerely be interested in knowing about it.
Now that having been said, it IS theoretically possible to use spray-cooling to increase the total pressure of an incoming air stream. This principle was first identified by Dr. A.H. Shapiro in his seminal paper “The Aerotheropressor- A Device for Improving the Performance of a Gas-Turbine Power Plant”, Transactions of the ASME, April, 1956. This concept exploits the cooling induced by an evaporating spray to increase the total pressure of a gas flow, effectively moving down the Rayliegh line. This concept was later exploited as part of a new engine concept, known as a “transition engine”, by myself in 1995 (“The Transition Engine: A Combined-Cycle Engine Concept for SSTO/Trans-atmospheric Vehicle Applications”, AIAA 95-2480).
The air ionization approach is also interesting and novel. I’ve done a little work using high voltage DC sources to create “ion winds”, and these do indeed make a bit of thrust. Its hard to induce a lot of air movement with this approach, but maybe there’s some clever geometries that can make this work. The power supply system and HV charging equipment is by no means a trivial problem with electric propulsion. I think this area could bear a lot of fruit in terms of mass reduction for power supply and conditioning systems, but that will take some serious and sustained development efforts to realize usuable gains there. FYI, for the electrical supply, the highest power density systems come from, yes that right, turboelectric generators. You burn some fuel, expand it through a turbine, spin a generator. These systems sometimes have an order of magnitude higher power (and energy) density than batteries, supercapcitors, RTGs, fuel cells, etc. ATRs already have the turbomachinery….hmm…..
A couple of points regarding the J58 and the SSME turbopumps. First, I believe the J58 turbine actually outputs around 160,000 shaft hp (http://www.hill.af.mil/library/factsheets/factsheet.asp?id=5786 ). As you probably already know, that the Thrust Specific Fuel Consumption (TSFC, and sometimes reduced to just SFC) is the flowrate in lbm per hr divided by the thrust produced (also, the inverse of the TSFC multiplied by 3600 gives the Specific Impulse. The above ref quotes a max fuel flow of 8000 gallons per hour, or about 14 lbm/sec, somewhat different that what you quoted at around 6.25 lbm/sec, but I don’t think this changes the argument.
The power required by a turbopump is proportional to both the fuel flow rate and the delta P that the pump creates. The required fuel pressure in the J58 is far lower than the 3500+ delta P required in the SSME. You could estimate the J58 fuel pump power required by multiplying the fuel flow rate by the delta P, then dividing by rho and 2. I show it being more like around 25 or 30 hp for an assumed delta P of about 300 psi. (I believe the J58 uses a fuel pump driven from an accessory gear off the main turbomachinery shaft. At least, that is a conventional arrangement).
Well, I hope I provided some useful feedback regarding your comments and ideas. In summary, I believe that Turbocompressors have evolved because they provide the trade between efficiency, pressure rise, delivered flowrate, and mass. And that’s why we use them. Despite being heavier, they offer some powerful logistical advantages. I guess that’s why I think the ATR offers some general advantages, but advantages that can only be truly appreciated when one looks at the overall system, where there are multiple, completing system requirements.
Nevertheless, other, novel approaches are well worth pursuing. It may be from some long-shot, high-risk approach that a viable and game-changing advantage comes from. And its guys like us who pursue these long shots, alone and ridiculed in our garages .
For an example of this, check the picture gallery on the plasma wind blog site for “other novel concepts”.
Good luck!
Thanks for the very informative response, Mr. Bossard. I do have an idea why the rated shaft horsepower for the J58 is wildly out of whack with the 1 horsepower per 2.5 pounds thrust estimate, even when you use the higher 32,500 lbs. thrust rating.
The 1 hp per 2.5 lb. of thrust estimate is rather close for the examples given on this page:
Convert Thrust to Horsepower.
…
“How much power does the 747’s Pratt & Whitney engine produce? As we discussed earlier, a static engine does no work no matter how much thrust it produces because it results in no motion. We must instead focus our attention on a plane that is in motion. For example, our 747 typically cruises around 600 mph (970 km/h). However, we are faced with a new problem because the plane does not necessarily need every bit of its static thrust to fly at that speed. In fact, static thrust is really an ideal maximum amount of thrust that an engine can produce in a test environment. As discussed in a previous question about thrust ratings, any jet engine will produce less thrust in actual use than the static value.
Furthermore, aircraft are equipped with throttles that allow a pilot to adjust the amount of thrust an engine produces. A good example is the SR-71 Blackbird equipped with Pratt & Whitney J58 turboramjets that produced a combined static thrust of 65,000 lb (289 kN). Even though the Blackbird could reach speeds in excess of Mach 3, however, it actually needed very little of this thrust in cruise flight. Most of the thrust was required to accelerate through the speed of sound, but once at Mach 3, the SR-71 engines were throttled back to only 30% or so.
“The conclusion of this explanation is that in order to determine the power a jet creates in flight, we need to know the exact amount of thrust necessary to fly at a particular speed. We typically know the static thrust rating of an engine or the airspeed of a plane during flight, but the problem is that we usually don’t know the amount of thrust that corresponds to a particular speed at a specific point in time. It is because of this disconnect that it is so difficult to calculate the power generated by the engines on a particular plane.
“Luckily, we do have access to data from a NASA report that does provide all the data we need to illustrate a sample case. The data is provided for a Boeing 747-200 cruising at Mach 0.9 at 40,000 ft (12,190 m). In this example, the aircraft’s engines produce 55,145 lb (245,295 N) of thrust, only a quarter of its rated static thrust, to cruise at a velocity of 871 ft/s (265 m/s). Using the equations provided above, we calculate the power generated by the 747 to be 87,325 hp (65,100 kW).
“The NASA data also includes a few other planes, so let’s compare the power generated by the subsonic 747 airliner to a supersonic fighter like the F-4 Phantom II. In this example, the F-4 cruises at Mach 1.8 at 55,000 ft (16,765 m). The aircraft’s two turbojet engines produce 11,560 lb (51,430 N) of thrust at its cruise speed of 1,742 ft/s (531 m/s). This combination of force and speed equates to a power of 36,620 hp (27,310 kW).”
http://www.aerospaceweb.org/question/propulsion/q0195.shtml
Note that these numbers are close to the 2.5 to 1 estimate. However, note also that these examples are basing it on how much speed the *entire aircraft* achieves. But clearly this would be dependent on a lot of factors of the general aircraft aside from the engine such as drag, weight, etc. And the article even suggests the horsepower would be zero in a static test environment. But clearly a static jet engine is still putting out alot of power in this case.
Another means of calculating the power of a jet engine would be by using the equation power = (thrust)x(velocity) mentioned on this page, but using for the velocity the exhaust velocity of the engine, not the aircraft’s speed.
Since the exhaust velocity of a turbojet is so high this probably explains the higher power rating for the J58.
This different method for calculating the jet engine power is probably closer to the one we need for calculating how much power we can bleed off for running a separate high performance, high pressure turbopump.
In regards to the ejector-ramjet analogue for creating high pressure through high velocity (supersonic) fuel injection, I’m inclined to agree it wouldn’t create enough pressure for the high volume of air we need.
However, I’m still optimistic about the possibility of achieving the compression of the large volume of air by using a wall of liquid propellant at very pressure, such as the 500 bar of the SSME liquid hydrogen turbopumps. I suggested before an annular ring opening for the liquid propellant inside the air intake. This may or may not be optimal. But for analyzing its feasibility let’s first just imagine this high pressure liquid being directed directly inwards into the intake from the front. It would be like a high pressure water hose being directed into the intake.
To insure the liquid and air just didn’t flow out the back you could constrict the cross-section so that the flow becomes somewhat choked. This would allow the pressure to build up before entering the combustion chamber.
You would need to allow the air also to enter the intake so you don’t want the hose to take up the entire intake opening. The idea however is to create a wall of liquid to compress the air. Perhaps a relative small gap between the hose outer diameter and the intake walls would allow most of the air to be so compressed.
As the air directly in front of the hose opening is moved inward, the incoming air from the sides *might* tend to fill in this space. However, the liquid is at such high pressure while the incoming air is at sea level pressure or less, I don’t know if this will actually take place. What might actually happen, and would actually be just as effective, is that the incoming air would be pushed against the sides of the intake wall by the liquid and be also thereby compressed.
The fuel output of the SSME turbopumps is much higher than the fuel need for the J58, as expected for the SSME producing so much more thrust than the J58. You might be able to create a much smaller version of the SSME turbopump for a J58 sized engine. (BTW, I’m aware the current fuel pumps for the J58 are so much less power than the SSME turbopumps. This undoubtedly is because the pressure in the combustion chamber in the J58 is so much less than in the SSME, and the required fuel flow rate is also so much smaller. However in my scenario I’m guessing you’ll need quite high pressure to create the “wall of liquid” to create the compression of such large amounts of incoming air. It would be so much the better if it could be accomplished using the relatively low pressure J58 fuel pumps!)
Creating the SSME turbopumps was a major technological challenge, however. I don’t know if they can simply be scaled down. Possibly you could instead use a larger version of the J58 engine and use the same SSME turbopump. Note that since you are dispensing with the compressors and turbines of the J58, just keeping the intake and combustion chamber, scaling this up might not be that big a challenge.
Alternatively, you could use several copies of the J58. Again because you made the engine lightweight by dispensing with compressors and turbines, this would still have quite good thrust to weight ratio, under the assumption the very high pressure fuel flow method can induce the usual compression ratio of the J58.
The feasibility of this method could be easily tested without using combustible fuels as the liquid. You could use water at the pressures seen with the SSME turbopumps of 500 bar. There are commercial water pumps that can put out these pressures. You could see if such high pressure water injected into a pipe with an increasingly constricted diameter caused the air to be compressed to pressures required for a turbojet engine, about 10 to 20 bar.
If this worked you might then try cryogenic liquid nitrogen to also see if the cooling effect you mentioned could improve the pressure increase even further.
Another key question that would need to be addressed is the mass/volume of liquid propellant being used to compress the large amounts of incoming air. From the the Wikipedia page on the J58 engine it uses 450 lbs/sec of air, 205 kg/sec. At an incoming sea level density of 1.2 kg/m^3, this is 170 m^3/sec of air as it first enters the intake.
Since as you say the fuel usage of the J58 is greater than I estimated, I’ll say the liquid hydrogen fuel usage might be larger as well, call it 4 kg/sec. Could this 4 kg of flowing liquid be used to compress 205 kg of gas (air)?
Said another way, the volume of this liquid hydrogen at a density of 72 kg/m^2 would be 1/18 m^3 every second. Could this be used to compress 170 m^3 of air?
Bob Clark
In regards to the method of converting the high pressure liquid hydrogen (500 bar) to a high velocity stream at ca. 1,100 m/s via the Bernoulli principle and using it in an analogous way to an ejector ramjet, your argument that this would be like a waterjet and would just cut through the air without compressing it (and possibly also through the intake walls) may indeed be correct.
Perhaps we could get it to be further like an ejector ramjet and therefore get the compression ejector ramjets get by vaporizing the liquid hydrogen at this high speed. We might be able to do this by heating just the outside of the pipe but more likely at this high speed we would need heated wires or concentrically arranged heated pipes within the larger pipe.
Bob Clark
I’m still wondering about that small amount of fuel at high pressure or high velocity being used to compress that large amount of incoming air.
Another possibility: use only a portion of the fuel in this way to compress its corresponding amount of air, then use the combusted fuel/air as a further ejector ramjet to compress the rest of the air and fuel. This would be like a two stage ejector ramjet. You could even extend this idea to several stages.
(BTW, in looking at some of the literature on “ejector ramjets” a.k.a “rocket ejectors” it appears the “ejector” is regarded as the part of the engine that draws in the air, not the rocket part as I was thinking. Puzzling, but I’ll try to be consistent with that usage.)
Say you wanted to use only half the fuel at the first stage. Let’s say we’re using liquid hydrogen with an air to hydrogen mass ratio of 40 to 1, about what you would get at stoich taking into account oxygen is about 1/5th of the total air.
Likely at this first stage you wouldn’t get much compression with this small amount of LH compressing the much larger amount of air. Then the thrust wouldn’t be especially good. But nevertheless after this first stage combusts, what you would get out would be high velocity exhaust at 40 times larger mass than before. Indeed it would now be at the same mass as the rest of the air that has to be combusted with the 2nd half of the fuel.
It wouldn’t necessarily have to be half the fuel you used at the first stage; it could be smaller or larger. Larger amounts though would mean for this part of the fuel you’re getting worse performance. So I’m inclined to believe, though am in no way certain, that smaller amounts would be more optimal for the first stage. You would need more than one combustion chamber using this method but since each combustion chamber is smaller hopefully the total mass would be the same.
If you used 3 stages, you could use 1/4th the fuel and air at the first stage. This would be combusted, at low performance, to compress the same amount of air and fuel, that is 1/4th, getting better performance, which combustion would be used to compress another 1/2 the air and fuel, getting even better performance. Again these proportions could be varied to see which gave the optimal thrust and Isp.
IF it turns out that you can get good final performance by using a significantly smaller amount at the first stage, say 1/4th, 1/8th, etc., then what you might do instead is just use a quite small turbine engine at this first stage. This would improve the performance at the first stage and the mass would still be significantly less than a normal turbojet since only the much smaller jet at the first stage has compressors and turbines.
Here are a couple examples of small turbojet engines used for manned aircraft:
MICROTURBO TRS 18-1
ENGINE SPECIFICATIONS.
http://www.bd-micro.com/FLS5J.HTM#ENGINE
Turbojet engine TJ 100.
http://www.pbsvb.cz/dlt_motor_tj100.php?lang=en
This last might also be something amateurs could test out. There are tiny jet engines meant for remote controlled model airplanes. Amateurs have also built examples of small ramjet engines using air from compressed air cylinders at the front to provide the compression. Then you could instead use one of these model airplane jet engines at the front of the ramjet to act analogously to an ejector ramjet.
Model airplane jet engines:
Advanced Micro Turbines.
http://www.amtjets.com/specs.html
Example of an amateur-made, compressed-gas ramjet:
Coffee Mug Ram Jet Engine.
http://www.youtube.com/watch?v=UUJP5Jh9_Bg
Bob Clark
John B., Bob, etc…
I’m not sure if anyone else knew this but PART of the reason the fuel numbers of the SR never seem ‘exactly’ right is because once it leaves the fuel tanks the mix was used (and along the way some inevitably gets lost I suppose) everyhere else in the fuselage/engine system FIRST! Directly from the tanks it was the hydralic fluid for the engine systems and flight controls, then would be used for cooling life support, the airframe, etc, and only at the very end of its journey would it be injected into the engines to burn!
John B. if you can look for an archive copy of or site-mirrors that saved some of the alt-acc.com information from when it was up and running. Before Plasmawind came along it was pretty much THE best place for air-breathing information in general and ramjet engines in particular on the web and a lot of the information was gathered no where else on the web.
Bob:
The Gluhareff Pressure jet engine, see:
http://www.tipjet.com/
http://www.tipjet.com/tech_data.htm
… uses high speed (supersonic flow) propane that has been super-heated to entrain air through a three-stage sonically tuned intake system with no moving parts with enough static thrust to get itself moving and once ram-air begins thrust increases to optimum levels with a smooth throttling control and fairly good SFC. The idea of using hydrogen for this same purpose (which is basically what’s being discussed as I understand it) is probably do-able though more likely with the fuel injected as a high speed gas after traveling through a series of heat exchangers, (hull, intake, combustion chamber, nozzles, the to injection) rather than being injected as a liquid.
Randy
Randy:
thanks for the reference link, always good to know more sources. You’re quite right. Because the fuel was required to perform multiple tasks, it was highly specialized, had lots of funky additives, etc. This contributed to it being hard to ignite, and why they went for TEB injection for ignition. While reliable, it was logistically complicated. (I’ve used TEA/TEB for igniting rocket engines. It works well, just don’t uncap the supply bottle. The green color comes for the boron compound).
Thanks for the info on the Gluhareff Pressure jet engine, Randy.
This is indeed quite analogous to what I was proposing for using high velocity injected fuel to induce pressure in the air. This method doesn’t even need high performance turbopumps for the fuel but gets it to high pressure by heating it from the heat produced by the engine.
(I still would like to see tried though the method of just using very high pressure fuel, not at high velocity, to compress the air by a “wall of liquid”.)
I was interested to see in the description of the Gluhareff jet that pressure was raised in several stages. I don’t believe though they are also using separate combustion burners at each stage.
You might be able to get even higher thrust by using more than one combustion chamber in the stages. The tech info on that Gluhareff engine only gave it a thrust/weight of about 4 – 5 to 1.
Bob Clark
Here are some other proposals different from the ejector ramjet idea, based rather on the idea of rotary rockets, that I sent to an amateur rocketry list.
Bob Clark
— On Sat, 8/1/09, Robert Clark wrote:
> From: Robert Clark
> Subject: Re: [AR] Another prize suggestion – Re: Hypersonic ‘WaveRider’ poised for test flight.
> To: “arocket”
> Date: Saturday, August 1, 2009, 6:59 AM
>
> There seem to be a million dozen different ways of doing
> the turbine-free air compression for jet engines. But it
> seems like all the jet engine and aircraft manufacturers are
> only focused on improving turbine engines. Why?
> Perhaps because turbine engines in their view are already
> good enough, at least on the measure we’re trying to improve
> on: thrust/weight. For while a rocket engine might have a
> T/W of 70:1 and a jet engine might only be 6:1, because of
> lift, where a typical jet aircraft might have a lift-drag
> ratio of 12:1, the T/W for the jet becomes effectively 72 to
> 1.
> What the jet engine professionals are very interested in
> is improving fuel efficiency since fuel costs are a big part
> of airlines profit margins. The methods suggested of doing
> away with compressors/turbines look like they will actually
> make fuel efficiency worse, such as the ejector ramjet for
> example, the exact opposite of what the jet engines pro’s
> are majorly focused on.
> But how about the rocket guys? They’re focused on getting
> to orbit and what happens at quite high Mach numbers. What
> happens at subsonic speed to low supersonic speed holds very
> little interest.
> So if this is going to be done, it looks like we’re going
> to have to be the guys to do it.
>
> And here’s another one of the million dozen different
> ways:
>
> I just saw this on Selenian Boondocks:
>
> Rotary PDE
> Nov 23rd, 2008 by johnhare
> guest blogger john hare
> http://selenianboondocks.com/2008/11/rotary-pde/
>
> Hare suggests using multiple combustion chambers within a
> rotating torus. He suggests this for rocket engines using
> pulse detonation propulsion. But the idea would also work
> for jet engines and you don’t need to use detonation
> for the method. Regular combustion would work.
> Detonation propulsion still is not well understood. Hare
> was suggesting the detonation wave to provide the
> compression, but you could just have the rotating torus
> generate the compression required for a jet engine by
> centrifugal force. The force to rotate the torus would be
> provided when each combustion chamber released its exhaust
> in turn.
> As Hare mentions, this idea for the rocket case is
> analogous to Gary Hudson’s rotary rocket, at least in the
> second incarnation where Hudson made the rotating
> engines be internal rather than on external rotors.
> Hudson then might be a good one to ask about its
> feasibility for the jet engine case.
>
John Hare discusses a variant of the Gary Hudson Rotary Rocket to get high thrust for the thrust provided by the rockets at the rotor tips here:
Roton Revisit.
Nov 5th, 2008 by johnhare
http://selenianboondocks.com/2008/11/roton-revisit/
In the comments section on this page, Wayne Gramlich also pointed to a page on an internal type of rotary rocket engine by Roger Gregory:
halfwaytoanywhere.com rotary rockets.
http://www.halfwaytoanywhere.com/
These are rocket engines. I wanted to investigate the case for airbreathing propulsion.
John Hare in responding about the Gregory rocket engine noted it might be dangerous to have the engine be rotating at ca. 600 m/s rim speed while at high pressure, 10,000 psi chamber pressure, and at the high temperature of combustion chambers.
This would be a problem in the airbreathing case as well. You might want to use the high velocity rotating torus just for compression and not have the combustion there. Firstly, how fast would you need to rotate it to get the ca. 10 to 20 bar compression typical of turbojets? Perhaps someone can give the equations for centrifugal compression of a gas given rotation speed and radius.
As a first guess I’ll estimate it by what happens with ramjets. Ramjets get good compression and thrust and Isp when the aircraft is flying in the range of Mach 2 to Mach 3, about 650 m/s to 1,000 m/s. I mentioned before a case of a ramjet able to get a thrust/weight ratio of 40 to 1 using hydrocarbon fuel. So we could get likewise good compression, thrust, and Isp if our rotating chamber encounters the air, when still as at start or when moving as when flying subsonically, at a rotating rim speed of 650 – 1,000 m/s. There are flywheel rotors that can rotate at these speeds, though I believe 1,000 m/s is near the edge of what is currently being done.
But this would be a problem if we made the high temperature combustion chamber be rotating at this speed. The highest rim speed flywheels use for example carbon fiber or synthetic fibers such as Kevlar to withstand the tensile stresses at these speeds. The chamber pressure when using air would not be nearly as high as the Gregory rocket case of 10,000 psi using liquid fuel and LOX. But still the materials used for fast flywheels could not withstand the temperatures of combustion.
Perhaps regenerative cooling could be used as for rockets. But the allowed temperature would be significantly lower for the flywheel materials compared to the metals used for rocket engines.
A few other possibilities. Use two torii, one for the combustion chamber, one for the compression chamber. The compression chamber torus may compress the air by encountering the incoming air at supersonic rim speed, as happens with ramjet compression, or just by using centrifugal compression. Using centrifugal compression, with the air introduced at the hub, might be preferred since you don’t have to deal with the shock waves or the heating from the low altitude, high density air suddenly encountering an inlet at supersonic relative speed. In any case, once the air is sufficiently compressed it would then be presented to the combustion chamber.
The combustion chamber would be rotating as I suggested before due to exhaust vented from nozzles around the toroidal chamber, as with Hero’s engine (like a lawn sprinkler.) Note in this scenario, unlike what I first proposed in a prior message, the compression is NOT provided by the rotation here. So you don’t need high rotational speed and can use higher temperature resistant metals. You convert this low speed rotation of the combustion chamber into high speed rotation of the compression chamber by principles of mechanical advantage, such as with gears, levers, etc.
Note you also have some flexibility with this combustion chamber. You might be able to get better efficiency and have lesser technical complexity if the combustion chamber does not rotate. Instead it could be a standard fixed combustion chamber but you could have some of the exhaust bled off to connect to a Hero’s engine, for lack of a better term. Or you could have instead use a heat eschanger to heat hydrogen using the heat only from the combustion chamber to thereby drive the Hero’e engine. Since hydrogen is of low molecular weight you wouldn’t need to get very high temperatures here to get high velocity exhaust from the nozzles of the Hero’s engine. As before the rotation of the Hero’s engine would be used to drive the rotation of the compressor torus.
There are several other variations. For instance instead of a toroidal compression chamber you could have a piston and cylinder arrangement driven by the Hero’s engine, as the piston is seen here driven by a flywheel:
Energy Storage II.
http://zebu.uoregon.edu/2001/ph162/l10.html
The jet engine would likely have to be pulsed in this case but we could emulate continous operationn by having two or more pressure cylinders operating sequentially.
As a variation on this idea we could use fan blades, cupped or flat, that drive the air for compression as suggested by John Hare here:
Cagejet Turborocket
Oct 17th, 2008 by johnhare
http://selenianboondocks.com/2008/10/cagejet-turborocket/
This would look analogous to a paddlewheel on a steamboat. The advantage of using this or of the piston and cylinder is that you wouldn’t need very high rotation or lateral speeds, just sufficient force to compress the large volume of air. Again this could be done by using methods of mechanical advantage, driven by the Hero’s engine.
Bob Clark
So, everyone, I hope these comments haven’t closed, because I just stumbled across them…
After wading through most of this stuff I figured I could just cut to brass tacks as I didn’t see my idea here already.
It’s pretty simple.
“TSTO on the cheap and easy!”:
Step 1 – Strip out interior of a common 747-8 transonic aircraft, including JP4. Inflate in-wing fuel bladders with helium to maintain wing structural integrity. Integrate interior of fuselage with LH2 cryotankage running the length of the aircraft.
Step 2 – Add ramscoops to outside of 747. Connect to a suitable heat exchanger “powered” by the onboard LH2. Add mechanical arrangements to fraction and separate the incoming air (we will use it in step 6).
Step 3 – Replace the 747’s 4 turbofans with 6 ramrockets optimised for subsonic flight – ramrockets to be powered by LH2 warmed by the onboard air liquefaction equipment.
Step 4 – Attach some kind of orbiter craft to top of 747, with orbiter to be powered by onboard aerospike engine(s) fueled by carrier-aircraft-made LOX and preloaded LCH4.
Step 5 – “Hop” from runway to air with orbiter (full of LCH4 only) attached to 747 (loaded full with LH2), proceed to cruising speed (Mach 0.8).
Step 6 – “Skip” oxidizer loading by cruising at speed with the air-liquifier on; the LOX is diverted to the orbiter tankage (loading it for an insertion burn with the LCH4) while the remaining liquid air fractions (nitrogen, etc.) are used as cold working mass to be injected into the ramrocket engines (I envision little aerospikes in them – the still-liquid air fraction(s) would be injected where the virtual “air spike” is supposed to sit in the inside of an annular rocket-nozzle/hot gas generator, rapidly expanding and improving outbound mass flow – incoming air should keep the “spike” tapered and make a good combustion zone).
Step 7 – “Jump” when the orbiter is full of LOX (and the astronauts of bagels – tee hee) and the 747 reaches a nice altitude like 40,000 ft; separate vehicles at subsonic speed, drop the 747 for the return journey to the runway and light the orbiter’s rockets for a burn to Mach 25 and Low Earth Orbit.
Inspiration: Shuttle Carrier Aircraft + NASA GTX + SpaceShipOne + LACE + Skylon + Black Horse + Pegasus(rocket)
Advantages: “Free lunch” (ha ha) ride for orbiter to 40,000ft and ~500mph, in-flight oxidizer loading utilizing LH2 fuel that will be burned anyway, subsonic vehicle separation (to conserve stability), kickass dead-simple (and relatively cheap) ramrocket engines on the carrier craft that will provide static thrust AND collect both working mass and oxidizer in-flight to improve performance, aerospike (overall most efficient!) rocket engine(s) on orbiter, everything is 100% re-usable.
Disadvantages: Heavy LH2 tankage, heavy orbiter (certainly!), heavy heat exchanger – all things that can be engineered lighter.
Oh, and a jumbo jet full of explosives… like the Shuttle External Tank, but with wings. Don’t be scared.
Am I missing anything?
Oh yeah – I like your turborocket ideas but seriously turbomachinery is heavy, expensive and has zillions more points of failure than a rocket (e.g. thin vanes pulling multiple G at ridiculous temperatures). A rocket with some duct wrapped around it (ramrocket!) is hardly more complex than a rocket alone and can provide static thrust also, plus this particular design will have several metric f*cktons of LH2 available to burn so using fuel-gulping engines means almost nothing. As for the orbiter – if all you can fit on it for cargo (minus pilots) is 1,000kg but you can reach LEO or even geosync each flight (with a fully reusable launch system) you return your bilion+ $ investment in the first system’s lifetime – anyone with something to get into space will buy your services.
BTW the 747 was a suggestion – you could also try an AN-225 or something else with sufficient lift capacity (ha!).
_d
P.S. – Criticize me! I am an ignorant youth. Also I want feedback.
JH: I found this in the filter and decided to paste them here and comment on them. It was posted by “Some Slob” I hope that’s just a screen name.:-)
So, everyone, I hope these comments haven’t closed, because I just stumbled across them…
After wading through most of this stuff I figured I could just cut to brass tacks as I didn’t see my idea here already.
It’s pretty simple.
JH: not really
“TSTO on the cheap and easy!”:
Step 1 – Strip out interior of a common 747-8 transonic aircraft, including JP4. Inflate in-wing fuel bladders with helium to maintain wing structural integrity. Integrate interior of fuselage with LH2 cryotankage running the length of the aircraft.
JH: You’ve just redesigned the wing and fusilage, that’s going to cost. A new vehicle might be cheaper.
Step 2 – Add ramscoops to outside of 747. Connect to a suitable heat exchanger “powered” by the onboard LH2. Add mechanical arrangements to fraction and separate the incoming air (we will use it in step 6).
JH: That’s a lot of modification and machinery thrown out there.
Step 3 – Replace the 747’s 4 turbofans with 6 ramrockets optimised for subsonic flight – ramrockets to be powered by LH2 warmed by the onboard air liquefaction equipment.
JH: Ramrockets are terrible for subsonic flight. Fuel consumption approaches the rocket alone. Plus you would have to develop them from scratch. Also, ramrockets need oxidizer in the rocket side.
Step 4 – Attach some kind of orbiter craft to top of 747, with orbiter to be powered by onboard aerospike engine(s) fueled by carrier-aircraft-made LOX and preloaded LCH4.
JH: See the airlaunch home page for the difficulties of top mounted orbiters.
Step 5 – “Hop” from runway to air with orbiter (full of LCH4 only) attached to 747 (loaded full with LH2), proceed to cruising speed (Mach 0.8).
JH: By hop, do you mean take off?
Step 6 – “Skip” oxidizer loading by cruising at speed with the air-liquifier on; the LOX is diverted to the orbiter tankage (loading it for an insertion burn with the LCH4) while the remaining liquid air fractions (nitrogen, etc.) are used as cold working mass to be injected into the ramrocket engines (I envision little aerospikes in them – the still-liquid air fraction(s) would be injected where the virtual “air spike” is supposed to sit in the inside of an annular rocket-nozzle/hot gas generator, rapidly expanding and improving outbound mass flow – incoming air should keep the “spike” tapered and make a good combustion zone).
JH:You have added a very complex subsonic engine cycle to a cutting edge air liquifier with no visible justification. This is turbofan country at cruise.
Step 7 – “Jump” when the orbiter is full of LOX (and the astronauts of bagels – tee hee) and the 747 reaches a nice altitude like 40,000 ft; separate vehicles at subsonic speed, drop the 747 for the return journey to the runway and light the orbiter’s rockets for a burn to Mach 25 and Low Earth Orbit.
JH: Air launches at 40,000 feet are available with almost unmodified aircraft.
Inspiration: Shuttle Carrier Aircraft + NASA GTX + SpaceShipOne + LACE + Skylon + Black Horse + Pegasus(rocket)
Advantages: “Free lunch” (ha ha) ride for orbiter to 40,000ft and ~500mph, in-flight oxidizer loading utilizing LH2 fuel that will be burned anyway, subsonic vehicle separation (to conserve stability), kickass dead-simple (and relatively cheap) ramrocket engines on the carrier craft that will provide static thrust AND collect both working mass and oxidizer in-flight to improve performance, aerospike (overall most efficient!) rocket engine(s) on orbiter, everything is 100% re-usable.
JH: See above on ramrocket engines.
Disadvantages: Heavy LH2 tankage, heavy orbiter (certainly!), heavy heat exchanger – all things that can be engineered lighter.
Oh, and a jumbo jet full of explosives… like the Shuttle External Tank, but with wings. Don’t be scared.
Am I missing anything?
JH: Simplicity. Calling a complex method simple doesn’t make it so. Complexity is only desirable in direct relationship to it’s advantages.
Oh yeah – I like your turborocket ideas but seriously turbomachinery is heavy, expensive and has zillions more points of failure than a rocket (e.g. thin vanes pulling multiple G at ridiculous temperatures). A rocket with some duct wrapped around it (ramrocket!) is hardly more complex than a rocket alone and can provide static thrust also, plus this particular design will have several metric f*cktons of LH2 available to burn so using fuel-gulping engines means almost nothing. As for the orbiter – if all you can fit on it for cargo (minus pilots) is 1,000kg but you can reach LEO or even geosync each flight (with a fully reusable launch system) you return your bilion+ $ investment in the first system’s lifetime – anyone with something to get into space will buy your services.
JH: Read up on turbines a bit. Fear of turbines seems to be a bad meme in rocket circles.
BTW the 747 was a suggestion – you could also try an AN-225 or something else with sufficient lift capacity (ha!).
_d
P.S. – Criticize me! I am an ignorant youth. Also I want feedback.
JH: Part of being an innovator is the requirment for a thick skin AND open ears. You will need to be able to hear critisizm without taking it personally, and the ability to filter the wheat from the chaff. Innovators without the thick skin give up too quickly. Innovators that don’t listen get killfiled as it’s a waste of time explaining to them. The middle of the road is a difficult path, but the most rewarding if you can hack it. Watch people shred my ideas to see what I mean. The ones ready for immediate use would be going to my friends in the industry direct instead of free on a blog.
So, everyone, I hope these comments haven’t closed, because I just stumbled across them…
After wading through most of this stuff I figured I could just cut to brass tacks as I didn’t see my idea here already.
JH: All these ideas have been tossed out individually from time to time. What kills most of them is the cost, risk, and complexity. If it is difficult to get one of them at a time funded, how hard do you think it wil be to get all of them at once done?
It’s pretty simple.
“TSTO on the cheap and easy!”:
John, have you looked at the ejector ramjet and supercharged ejector ramjet concepts studied by Marquardt in the late sixties? They too are concepts capable of operation between Mach 0 and Mach 5. The supercharged ejector ramjet can also be used in ducted fan mode, acting as a very efficient low speed engine during descent and landing. This gives you crossrange, which is nice.
Another random idea: what about peroxide and subcooled methanol? Methanol is a good coolant and so is peroxide. Peroxide is very dense and methanol is fairly dense. The melting point of methanol is -97C. During airbreathing such an oxygen containing fuel might benefit less from atmospheric oxygen, but because of the more balanced O/F ratio it could allow for better thrust.
Martin,
It is my opinion that the turborocket based concepts are far superior to the ejector series below mach 2. Above that, you really want to be out of the atmosphere soon enough that the advantage range is quite small. Turbine based systems are actually easier in this field than iaircraft use since we can trade efficiency for simplicity and thrust/weight. From a standing start, I believe that a turbine based system could be fielded faster and cheaper than the ejector series if you factor in testing and vehicle integration.
@ j hare -
I suppose my post makes more sense if you consider the question to have been “how would *I* get to space” rather than “how would *we* get to space”.
Let’s say you’re a country who’s just gotten out of 3rd-world debt and wants to move toward space. Ejector ramjets offer shorter and less expensive development/testing/manufacturing lifecycles than turbojets considering you don’t own any turbojet IP or have any commercial giants (Boeing, Ariane, etc.) clamoring to set up shop in your borders. The fact that turborockets also have yet to ever fly will make them even more costly than ramrockets from scratch – if you’re a brand-new space explorer with zero experience in advanced avionic technology trying to integrate someone else’s proprietary engine without melting it OR violating patents… you see what I mean? Growing your own IP fruit tree will be easier than trying to embed someone else’s turbofan design in it as it grows. As for your reply:
1 – The whole point of modding a 747 is that they’re everywhere – you only have to ADD to them w/out redesigning the airframe to make them available for air launch – brand-new spacepower nations like Mexico would be able to afford something that easy.
2 – As above – all the “modification and machinery” added is something that has been done for years. The Shuttle Carrier has been extensively modified but remains un-redesigned – it’s still a 747. Simply bolting equipment onto an existing structure that is never expected to pass Mach 1 is a time-honored tradition in avionics; see Boeing 737 AEW&C, etc.
3 – You got me – I forgot about the ramrocket oxidizer. You could siphon it in-flight from the LOX-maker or carry another internal tank. The point I was trying to get across is that while ramrockets would SUCK for subsonic flight you’d have massive amounts of fuel (a 747’s-fuselage-worth) just waiting to be burned anyway; the subsonic portion of the flight with the carrier craft DOESN’T MATTER when considering efficiency – it’s not GOING ANYWHERE important, just to the landing runway to be refueled. If you’re Guiana and you’ve just built one of these carrier/orbiters France will start paying through the nose for flights – it doesn’t matter how efficient they are if you can leave the gravity well repeatably and reusably. Once the technology starts rolling then you can kaizen it – it’ll get more efficient over time when you have the money to improve it, plus you’ll OWN it (specs & all) outright.
4 – The video of Shuttle Enterprise holding stable and level while the SCA drops off (check wikipedia) sounds a lot louder than all your naysaying about top-mounted air launch. I believe what I see and what I see is a heavy, chunky b*tch of an aircraft with horrible subsonic aerodynamics and poor fuel efficiency just hanging there in glide while the carrier drops away. If NASA can do it… everyone else can do it cheaper.
5&6 – It’s cutting edge and complex and added to the subsonic engine cycle because carrying all that hydrogen onboard would be a waste if it wasn’t used to fly the plane. Carrying JP4 (and you’d have to with turbojets/turborockets – H2 embrittlement would kill a turbofan’s alloys, necessitating replacement) would defeat the purpose of stuffing the plane fuselage w/ fuel, which would negate the air liquefier, which would kill in-flight oxidizer loading – which, you have to admit, is a pretty sharp idea.
7 – Air launches are available with completely unmodified aircraft… if you jump from them (ha ha). Seriously though the in-flight oxidizer loading is why I say modify this stuff – simply carrying the orbiter on a (mostly) unmodified jet and tanking it up in-flight with a third aircraft would be troublesome with cryoliquids. Also you’d have another big-ass subsonic craft wanging around your airspace while you try to get something done. In retrospect it’s probably even cheaper overall than in-house development of anything but I am a slave to aesthetics… 1+1 = orbit, no remainder. Also the ramrocket-powered-carrier-craft would be MY next logical step in development after a 3-craft stage anyway, because I am young and have plenty of time for that sort of thing, plus I could live off the IP when I get old.
So you poked really nice holes in my theory, which I wanted. I hope you can get over your own close-mindedness about turbofans – just because everyone is circumcised (i.e. using turbofans) does not make it the most comfortable thing in the world. I was trying to consider what might present the (currently) most easily available AND most easily upgradeable solution to get to orbit for countries “just starting out” in the space race; e.g. what would *I* do if I suddenly had a nation-sized budget for spaceflight and wanted to keep doing it all the time. Developing ramrockets early on would make for savings and simplicity down the road, plus IP you could sell to other nations emerging from terran isolation. I admit I don’t like turbojets but I’d totally consider using them if they weren’t so damn complex mechanically AND intellectually – I could certainly buy an efficient, well-specced P&W turbofan on the cheap but I’d ultimately have to go back to P&W for spare parts, repairs, upgrades, etc. and they’d never tell me exactly how to build my own – that bothers me.
Overall I’m sure we can agree that nobody is trying hard enough to get us off this rock…
_d
I suppose my post makes more sense if you consider the question to have been “how would *I* get to space” rather than “how would *we* get to space”.
Let’s say you’re a country who’s just gotten out of 3rd-world debt and wants to move toward space. Ejector ramjets offer shorter and less expensive development/testing/manufacturing lifecycles than turbojets considering you don’t own any turbojet IP or have any commercial giants (Boeing, Ariane, etc.) clamoring to set up shop in your borders. The fact that turborockets also have yet to ever fly will make them even more costly than ramrockets from scratch – if you’re a brand-new space explorer with zero experience in advanced avionic technology trying to integrate someone else’s proprietary engine without melting it OR violating patents… you see what I mean? Growing your own IP fruit tree will be easier than trying to embed someone else’s turbofan design in it as it grows. As for your reply:
JH: Ramrockets sound like dead easy solutions on the first pass. On the second pass some of the problems start showing up. They would have to be developed from scratch in your scenerio. Straight rockets would be faster and cheaper. You really don’t want to fly them subsonic for any length of time, they use over ten times the fuel of a turbofan in that regime. For a straight climb to 50,000 feet from the runway with a massively overweight bird, maybe. For cruise while playing LOXmaker, no way as you will run out of aircraft fuel well before you get your spacecraft LOXed. In your scenerio, the Microcosm or Beal route would make more sense on the KISS principle.
1 – The whole point of modding a 747 is that they’re everywhere – you only have to ADD to them w/out redesigning the airframe to make them available for air launch – brand-new spacepower nations like Mexico would be able to afford something that easy.
JH: In your original, you suggested massive renovations.
2 – As above – all the “modification and machinery” added is something that has been done for years. The Shuttle Carrier has been extensively modified but remains un-redesigned – it’s still a 747. Simply bolting equipment onto an existing structure that is never expected to pass Mach 1 is a time-honored tradition in avionics; see Boeing 737 AEW&C, etc.
JH: The machinery you suggest has never been demonstrated in flightweight hardware.
3 – You got me – I forgot about the ramrocket oxidizer. You could siphon it in-flight from the LOX-maker or carry another internal tank. The point I was trying to get across is that while ramrockets would SUCK for subsonic flight you’d have massive amounts of fuel (a 747’s-fuselage-worth) just waiting to be burned anyway; the subsonic portion of the flight with the carrier craft DOESN’T MATTER when considering efficiency – it’s not GOING ANYWHERE important, just to the landing runway to be refueled. If you’re Guiana and you’ve just built one of these carrier/orbiters France will start paying through the nose for flights – it doesn’t matter how efficient they are if you can leave the gravity well repeatably and reusably. Once the technology starts rolling then you can kaizen it – it’ll get more efficient over time when you have the money to improve it, plus you’ll OWN it (specs & all) outright.
JH: The fuel use of the carrier aircraft does matter when there is so much of it required that it eats away the payload capacity.
4 – The video of Shuttle Enterprise holding stable and level while the SCA drops off (check wikipedia) sounds a lot louder than all your naysaying about top-mounted air launch. I believe what I see and what I see is a heavy, chunky b*tch of an aircraft with horrible subsonic aerodynamics and poor fuel efficiency just hanging there in glide while the carrier drops away. If NASA can do it… everyone else can do it cheaper.
JH: It can be done yes, but you can’t handwave away the issues involved.
5&6 – It’s cutting edge and complex and added to the subsonic engine cycle because carrying all that hydrogen onboard would be a waste if it wasn’t used to fly the plane. Carrying JP4 (and you’d have to with turbojets/turborockets – H2 embrittlement would kill a turbofan’s alloys, necessitating replacement) would defeat the purpose of stuffing the plane fuselage w/ fuel, which would negate the air liquefier, which would kill in-flight oxidizer loading – which, you have to admit, is a pretty sharp idea.
JH: The carrier aircraft will have to carry the whole mass of the orbiter oxydizer by the time of launch. IMO, it would be more efficient to launch with a full load of LOX as opposed to making it in flight unless there is a long cruise time built into the flight plan. That way your high altitude launch would be with a fully fueled orbiter from an aircraft with just enough fuel to get home on empty and no massive liquification machinery to lift.
7 – Air launches are available with completely unmodified aircraft… if you jump from them (ha ha). Seriously though the in-flight oxidizer loading is why I say modify this stuff – simply carrying the orbiter on a (mostly) unmodified jet and tanking it up in-flight with a third aircraft would be troublesome with cryoliquids. Also you’d have another big-ass subsonic craft wanging around your airspace while you try to get something done. In retrospect it’s probably even cheaper overall than in-house development of anything but I am a slave to aesthetics… 1+1 = orbit, no remainder. Also the ramrocket-powered-carrier-craft would be MY next logical step in development after a 3-craft stage anyway, because I am young and have plenty of time for that sort of thing, plus I could live off the IP when I get old.
So you poked really nice holes in my theory, which I wanted. I hope you can get over your own close-mindedness about turbofans – just because everyone is circumcised (i.e. using turbofans) does not make it the most comfortable thing in the world. I was trying to consider what might present the (currently) most easily available AND most easily upgradeable solution to get to orbit for countries “just starting out” in the space race; e.g. what would *I* do if I suddenly had a nation-sized budget for spaceflight and wanted to keep doing it all the time. Developing ramrockets early on would make for savings and simplicity down the road, plus IP you could sell to other nations emerging from terran isolation. I admit I don’t like turbojets but I’d totally consider using them if they weren’t so damn complex mechanically AND intellectually – I could certainly buy an efficient, well-specced P&W turbofan on the cheap but I’d ultimately have to go back to P&W for spare parts, repairs, upgrades, etc. and they’d never tell me exactly how to build my own – that bothers me.
JH: Don’t make the mistake of believing that anyone that disagrees with you has not researched the problem. Used turbofans will be cheaper than ramrockets for the forseeable future. You can’t sell IP (whatever you mean by that) if it is worthless to the potential customers.
Overall I’m sure we can agree that nobody is trying hard enough to get us off this rock…
Some people are trying hard enough, it’s just that the ones that are haven’t been in business for very long, and don’t have the bottomless pockets of the public purse to draw from. The previous government efforts have been like hitting someone with a stick, it hurts but doesn’t stop them, and the bruise will heal. The new space profit based efforts are like hitting them with a virus, they don’t notice until it spreads all over and stops them for the count.
_d
This makes it a good choice as an endo-atomospheric intereceptor engine, and as an airbreathing component to a space launch system. Interestingly enough and almost as an aside, I would note that, in principle the GG-cycle ATR will in fact be able to produce thrust even in a vacuum, although at a far-off-design operating condition.