Feeding the Turborocket

guest blogger john hare

Air breathing engines have many problems competing with rockets on acceleration missions. High on the list is the inlet that slows the incoming air and uses that kinetic energy to compress the air before sending it on to the engine. Air breathing engines really really don’t like supersonic, or even very high subsonic air at the engine face. The inlet efficiency is crucial to any high speed air breathing vehicle, and it weighs as much as the engine in many cases. Listed efficiencies for tested units seem to be from 90+% subsonic to <40% at higher mach numbers that are still well below scamjet range. Inlet designers seem to use the term pressure recovery to describe efficiency.

The inlet efficiency issue seems to be not understood to many ABE fans. I have seen suggestions that the boundary layer just be used in the engine to avoid the whole issue. At mach 2, a given inlet might compress air by a factor of four. If the engine behind this inlet is eating 1,000 cubic feet of air per second, then sucking in the boundary layer air would give the engine one fourth of the mass per second as using the inlet. At sea level, that is roughly 70 lbs/sec vs 280 lbs/sec for the same engine. A second issue is that increasing chamber pressure by a factor of four does very good things to the Isp, like doubling it or so. There can easily be a factor of eight  thrust difference with a good inlet vs no inlet at mach 2.

 As the mach numbers climb, inlet efficiency becomes more critical yet. When potential compression ratio is 100, the difference between a 60% and 30% efficient inlet become the difference between acceleration and deceleration in a vehicle that typically has little excess thrust to begin with. This is where ramjets can fail. They need an inlet with compression ramps that adjust to the mach number in the best possible manner without busting the mass budget. Though turborockets are somewhat less sensitive, they are still dependant on a clean, consistent, well compressed airflow if they are going to perform their mission.

Inlet compression at supersonic speeds consists of having surfaces that create controlled shockwaves that compress the incoming air. Every mach number, and fraction of one, have a best angle for the compression shock, and differing numbers of shocks increasing with mach number. At mach 1.3, a single normal shock is sufficient to slow the air  to something usable while at mach 2, you need at least one oblique shock before a normal shock leading into the diffuser that settles the air before the engine. There is a limit to how much each shock slows and compresses the airflow. As the mach numbers climb, you want more oblique shocks of smaller angle to the airflow to recover as much pressure as possible. By mach 5, you want a dozen or more if you can get them, each at a very few degrees to the airflow that was turned by the shock before it. Isentropic ramps are a smooth curve that theoretically gives an infinite number of very small shocks for the best possible performance. The problem is that the best angle of each mini shock changes with mach number which prevents just building a simple solid inlet and using the inside of it for something useful like a fuel tank. People in the business and some references refer to the difference between drag and thrust at high mach as a very small difference in some very large numbers.

The normal method of dealing with varying mach numbers is to have ramps that shift angles scheduled with the mach number to give as much efficiency as possible. Most of these ramps are curved in one direction and flat in the direction normal to the airflow. They are not perfect for every mach number, just the best compromise for the expected use of the vehicle. The structure to keep these moving ramps intact and in place in high mach conditions, and the mechanisms to handle them, can get heavy fast if you are after an acceleration mission.

Then there is the boundary layer that builds up in the adverse pressure gradient of the inlet ramps. If it builds up too much in the wrong manner, it can trip the whole shock system and virtually eliminate pressure recovery. Either the layer must be bled off and dumped or reenergized toward the pressures of the rest of the stream. Elaborate systems are employed to bleed the layers to avoid turbulence, but not so much that good compressed air is being dumped. Figuring out the best methods of handling this boundary layer problem is a job for a full time team.

Then there is the temperature problem. As air is compressed, it heats up. By mach 4 or so, turbomachinery becomes a liability just on the heat issues alone. Turbine inlet temperature become excessive on turbojets even with low compression ratios about here, while the turborocket can get a couple of more mach numbers because the turbine doesn’t see the compressed air. Precooling is the answer most suggested for this mach range, and is a whole nother engineering field again. Physical precoolers can get fussy about excessive moisture in the air freezing and foreign object damage is like a rock through your car radiator. Not to mention they get heavy fast. Mass injection is another solution with possibilities. Liquid oxygen or water is injected into the stream in front of the compressor to cool the incoming airflow to something useful to the engine. At some point though, the inlet air is so hot that cooling mass flow resembles rocket mass flow for the same mach number, and the rocket isn’t dealing with all that heat and other problems of the ABE.

While I don’t explain everything well, and probably got a thing or three wrong so far, I think it is obvious why rockets are the preferred acceleration engine. All the problems above are with engine types that quit somewhere below mach 6 when a rocket is just getting started. Anybody advocating air breathing engines needs to be honest about the issues they have for spaceflight. Many are not, handwaving problems they don’t understand, that the rocketmen they talk to do know about. When ignorant of the field they are advocating, they destroy their own credibility, and damage that of sincere workers. At the moment though, we can run these things past somebody that is honest, competent and open on the subject, John Bossard at http://plasmawind.typepad.com

This is one more of my tech thoughts, this time on solving some of the issues above in a manner that works for spaceflight.

 A turbine spins to drive a compressor with a combustor between them to heat and energize the mass flow in a turbojet. An inverse cycle uses the incoming airflow to drive a turbine that powers a compressor before fuel is added. As far as I can tell, the cycle has never been demonstrated to anybodies satisfaction, they have to put combustors in there somewhere.

I propose designing a turbine wheel to operate as the inlet with the rpms increasing with flight speed to vary the angle that the incoming air sees.The supersonic air hitting the turbine wheel is turned though an isentropic curve with the initial angle determined by the relative mach numbers of the turbine and airflow. This eliminates moving ramps to schedule inlet angles. The centrifugal action of the blade on the air re energizes the boundary layer to similar values as the rest of the incoming stream. This eliminates the complex bleed systems. To keep the turbine from over speeding, it is connected to a compressor stage with a stator group between them. This both controls the turbine rpm and improves the pressure recovery of the total inlet.

turboinlet

This idea would take a serious research effort to reach hardware if it is feasible in the first place. If it is feasible, it will make for a much lighter inlet, with good efficiency across several mach numbers, and a much shorter operational length than the current systems. It could make the difference between feasible and not for a marginally capable engine system.

If that system will solve the shock angle and bleed air problems, it still does nothing about the severe thermal environment at high mach numbers. I suggest regenerative cooling of the rotating inlet by injecting large amounts of very cold water.  The water at less than 40 degrees f flows through the inlet turbine, stator, and compressor blades and out into the inlet airflow. Inject so much cold water that the temperature of the airflow drops under 100 degrees f even behind the inlet compressor. Then recover the water in a vortex separator and send it to a cooler before reusing the same water.

 precooler

In the lower atmosphere at the lower speeds, the cooling water can even increase in mass by collecting moisture from the air and drying it out with the cold air. There are no heavy, complex, and finicky physical precoolers, and the mass injection losses are quite moderate until  late in the flight when expending excess mass is acceptable. Relatively small amounts of hydrogen will cool large amounts of water with splorge cooling before being used to drive the turborocket. Late in the flight when lower turborocket rpms are acceptable, no LOX is needed for the turborocket with the gaseous hydrogen driving the turbine after cooling the water.

If something like this allows a turborocket first stage to accelerate an upper stage to mach 7 and 120,000+ feet, then even a dense propellant upper will have a mass ratio to orbit of well under 6. Economics of the market could decide on the viability of a system like this, I think favorably.

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19 Responses to Feeding the Turborocket

  1. Tim says:

    Hi John,

    To be clear, the inlet turbine is driven only by the incoming air, and no power source?

    I have to wonder if injecting the cooling water into the inlet airflow is worth it. The vortex separator looks complicated, and would potentially allow air to enter the splorge cooling tank where it could mix with the fuel. Collecting the cooling water in manifolds at the end of the inlet turbine and inlet compressor blades would create a completely closed system. Would you require a pump for the water, or would the turbine/compressor act as a centrifugal pump?

    One final thing, I assume “LOX as required” means that there would be airflow through the gas generator, as in your previous turborocket post?

  2. john hare says:

    Incoming supersonic air is a massive drag source for an airbreather. This is just turning some of that drag into more useful work.

    Water injection is a mature tech for increasing thrust of jet engines, though not through the blades as I suggest here. The vortex separator is based on a Russian paper concerning in flight LOX collection. A small amount of air in the fuel may be a problem, though I think the solutions are not that difficult. A small secondary seperator at the worst.

    There would need to be a pump for the hot water and the hydrogen before they enter the splorge cooling chamber. Otherwise the hydrogen would have to be compressed as a gas for the turborocket.

    This one is drawn as a true turborocket with the turbine driven by a rich H2/O2 combustion. As flight speed increases and more hydrogen is heated in the splorge exchanger, less O2 is needed to drive the turbine as it becomes an expander cycle. Some really crazy Isp becomes possible on the high end. Even subcooled methane should bust 4,000 at high thrust if everything went right.

  3. jsuros says:

    Two questions:

    1) How would the mass of water used and lost by this system compare to the mass of propellants?

    2) Could the hydrogen expansion as it cooled the water be used to pump the water back into the inlet assembly? Using several cooling chambers in series as piston-less pumps maybe?

  4. john hare says:

    1. Until flight speed gets fairly high, very little water will be expended at all, might even gain a little down low. From mach 3 or so on, increasing amounts of water will be expended. It would be nice to use an oxydizer for this application if the right one existed and it didn’t force a physical heat exchanger to be built. Using fuel instead creates a major fire hazzard.

    2. It depends on the pumped pressure of the hydrogen (or methane) but it sounds feasible. Pump pressure requirements will be fairly low as the turboinlet and compressor blades are going to be pumping also.

  5. jsuros says:

    How about using hydrogen peroxide as the cooling fluid?

  6. johnhare says:

    I’ve never even thought about peroxide for this. It would force a physicall heat exchanger, though if it could be liquid-liquid, it might not be too bad. Perhaps it could be splorge cooled with the LOX going to the gas generator. My first major concern would be thermal decomposition among the rotating machinery. “What is the liquid range of peroxide under these conditions?” is another concern.

  7. PeterH says:

    Before a rocket or jet engine can use propellant, the propellant must be brought to flight speed. (scramjets cheat a little here.) An air breather can extract energy from the incoming airstream. If energy is extracted with enough energy The air breather comes out a winner despite the lower chemical energy of the propellant. But as you explain, that’s very difficult.

    Sounds like what you’re proposing is to convert the incoming air kinetic energy to shaft energy, then use the shaft energy to compress the air. As you say, also not an easy proposal. Contrast the conventional inlet that attempts to directly convert the kinetic energy to adiabatic compression. Assuming the supersonic turbine works well, there might be other ways to use that shaft energy in the engine.

    As far as inlet cooling, one of the older and bolder proposals I’ve seen is to inject the fuel into the airstream in the inlet, upstream of the combustor. No heat exchanger needed, fuel is well mixed. Drawback is obvious.

  8. Mike Lorrey says:

    This is one of those situations where an insistence on perfect becomes the enemy of the good enough.

    Keep in mind that, for instance, the Falcon 1 first stage separates at under mach 6. Most rockets burn 2/3 to 3/4 of their propellant before they reach mach 6, most of the mass of which is LOX, which is of course eliminated from the mass budget of an air breather for that phase of flight.

    So really, an air breathing first stage makes a lot of sense not only on a fuel mass budgetary standpoint, but by making it an air breather, you are already on your way to making it a recoverable, reusable first stage, with all the potential cost benefits that come from that too.

    A mach 6 first stage would generally be a chassis cross-applicable to other HST needs. The rocket propelled second stage may or may not be reusable, generally depending on whether the payload is mass or men.

  9. Randy Campbell says:

    John H:
    There is a newer-type inlet known as the “Inward-Turning” inlet, some illustrations of which can be found here”
    http://www.astrox.com/SIDEBrochure.pdf

    A patent can be found here (I have a free account at freepatentsonline, I will follow with the patent number and title so others can use thier favorite patent look-up engine :)
    http://www.freepatentsonline.com/6164596.pdf

    Refernces for search are: US Patent 6164596 – Designs of and methodology for inward or outward, and partially inward or outward turning flow hypersonic air-breathing and rocket-based-combined-cycle vehicles

    Randy

  10. john hare says:

    Randy,
    I have the inward turning in my dead tree references. It is a favorite for missle use if we are talking about the same inlet. I spent a few bucks on the two textbooks available from AIAA on the subject. English authors have me wanting to call them intakes though.

    Mike,
    From a cost standpoint, the rocket vehicle will win if there is no cruise component to the flight. As will be pointed out again here by somebody, LOX is cheap. Air is free, but the convenience store charges more for it than gasoline.

    The value of an ABE is in flexibility given by a cruise phase such as fly back, fly forward, and loiter time if required.

  11. John Bossard says:

    John H:
    You’ve provided a very good post here. Thank you also for the “tip o’ the hat”, you do me great honor, sir.
    I will be providing some additional comments to this post, but I’m a bit behind right now. I’ll try to provide some comments within 24 hours.

  12. Randy Campbell says:

    John H. wrote:
    >I have the inward turning in my dead tree references. It is a
    >favorite for missle use if we are talking about the same inlet.
    >I spent a few bucks on the two textbooks available from AIAA
    >on the subject. English authors have me wanting to call them
    >intakes though.

    We “may” be talking about the same inlets though that’s a point the literature and web information can be confusing on. It’s one of the reasons I included the link to the ASTROX company:
    http://www.astrox.com/

    More specifically I should have included a link to their airbreathers image page:
    http://www.astrox.com/AirBreathers.aspx
    and the HySide brochure page (page 1)
    http://www.astrox.com/SIDEBrochure.pdf

    As you’ll note the “inward-turning” design in the illustrations looks similar to a “C” on its side (with the opening down) and it IS open all most all the way back to the combustor. It opens again directly after the combustor and remains so all the way to the rear of the vehicle. Though most ‘inward-turning’ inlets shown on vehicles show them fully enclosed from inlet to exhaust (as depicted in most of the Falcon, or HTV-3 images) this “new” style has emerged within the past 4 or so years to be the style of choice in most studies due to lowered mechanical complexity, (it does not require adjustable ramps as the “slot” seems to allow for self adjusment of inlet air at the desired Mach) less heating as the slot automatically ‘dumps’ excess air without needing mechanical dump vents as well as because the inlet is more streamlined and the overall vehicle aerodynamics are far better than either the “outward” or 2D installations previously studied.

    Though most of the material I’ve read on this inlet are predicated on the use of Rocket Based Combined Cycle engine (RBCC) due to the percived mass-penalty of turbine engines I believe that coupled with the ATR this design could achieve some serious mass saveings over a ‘standard’ Turbine Based Combined Cycle engine system.

    Randy

  13. john hare says:

    I looked at the link before but just couldn’t see enough detail. Is there a good paper on the concept?

  14. Randy Campbell says:

    John H. wrote:
    >I looked at the link before but just couldn’t see enough detail.
    >Is there a good paper on the concept?

    Several actually, though mostly they consist of Thesis written using the ASTROX programs for parametric studies and comparisions of Air-Breathing and Rocket launch vehicles. If you can’t link to these or see them let me know. I think I have copies I can email.
    http://oai.dtic.mil/oai/oai?verb=getRecord&metadataPrefix=html&identifier=ADA451531
    (Click the “Handle/Proxy” link to get to the paper or try this url)
    http://www.dtic.mil/cgi-bin/GetTRDoc?AD=ADA451531&Location=U2&doc=GetTRDoc.pdf

    This one you probably can’t get but:
    https://www.afresearch.org/skins/rims/q_mod_be0e99f3-fc56-4ccb-8dfe-670c0822a153/q_act_downloadpaper/q_obj_2c119f70-791d-4fd0-aadd-05908765ac94/display.aspx?rs=enginespage

    http://www.dtic.mil/cgi-bin/GetTRDoc?AD=ADA451531&Location=U2&doc=GetTRDoc.pdf

    The appendix sections on this one provide a lot of interesting information not only on the Inward-Turning design but others as well:
    http://www.docstoc.com/docs/995500/Reusable-Military-Launch-Systems-(RMLS)
    The original document can be found here:
    http://www.dtic.mil/cgi-bin/GetTRDoc?AD=ADA477059&Location=U2&doc=GetTRDoc.pdf

    article on ASTROX and U-of-M work:
    http://www.mtech.umd.edu/news/press_releases/mips_astrox.html

    If these don’t help let me know and I’ll see what else I can find

    Randy

  15. John Bossard says:

    A great post, John. As you discuss in your post, inlets for airbreathing engines, especially for high speed flight, are indeed a central issue, and pose some of the most difficult technological challenges for efficient airbreather operation. The issues only get more difficult as the flight speeds increase.
    Inlets play a central role in recovering total pressure from the free stream conditions, and as my old turbomachinery professor told us, “The total pressure through the engine (airbreather) is a fundamental performance parameter”, or words to that effect. If you can’t recover the total pressure, you can’t make thrust, no matter how hot you make the gas.
    There are a number of different definitions or conventions that one can use to quantify the performance of the inlet. In modeling I’ve done, we often use simply an empirically derived pressure-recovery schedule, which specifies the ratio of recovered stagnation pressure to freestream stagnation pressure as a function of free stream mach number. An analytic function that is reasonable and robust is the use of the so-called kinetic energy efficiency, as given by Kerrebrock (Aircraft Engines and Gas Turbines, J.L. Kerrebrock. ISBN-10:0-262-11162-4).
    The challenges you have enumerated for high-speed compressible flows are very real, can be somewhat counter-intuitive, and are not always easy visualize. A good example is when an inlet experiences an “unstart”. Without variable geometry, the inlet cannot be “restarted” without decelerating to a lower speed such that the shock can be “swallowed”.
    Isentropic inlets have the best theoretical pressure recovery but, for a fixed geometry, have only one design-point operating condition. At other flight speeds, that are operating at off-design conditions. Variable geometry isentropic inlets have been experimentally evaluated, but none have gone into service to my knowledge.
    I’m glad you mentioned the fact that, for airbreathing engines, the net thrust developed is essentially the difference between the so-called gross thrust, (mdot_fuel + mdot_air) x V exhaust, and the ram drag, mdot_air x V flight, (I’m assuming exit pressure force is negligible, sorry for the ASC equations, also see http://www.grc.nasa.gov/WWW/K-12/airplane/turbab.html ). At high flight speeds, the net thrust can indeed be a small difference between two large numbers. And this fact has a practical consequence. Small disturbances, such as might result from atmospheric conditions or vehicle attitude, can result in the net thrust going to zero or negative. This puts a practical upper limit on airbreathing propulsion, a limitation not imposed on rocket propulsion. And this is another reason why you need some performance margin in your airbreathing propulsion system.
    I’d like to comment in particular about inlet precooling, and also about the so-called “Mass Injection Precooling for compressors” or MIPCC. A number of engine concepts over the years have suggested various forms of inlet pre-cooling. Pre-cooling is helpful in a number of important ways. One of the most significant is that it increases the density of the incoming air. For Turbocompressors, the compression efficiency goes up as the air density increases. Some of these concepts have proposed using heat exchangers within the inlet duct, and running cryogenic liquids through these exchangers. A physical limitation to this approach is that, based on the Reynold’s Analogy, the total pressure gains from Rayleigh cooling will always be less than the total pressure losses associated with friction, i..e. Fanno flow. Thus, although inlet precooling using heat exchangers does increase the air density, it still results in a net loss on total pressure increase. Spray cooling of the inlet flow, however, is a different story. As A.H. Shapiro described in his paper “The Aerothermopressor- A device for improving the performance of turbomachinery” (Trans. of the ASME, April, 1956) the injection of liquid droplets into an inlet flow can, depending on how its done, actually result in a net rise in total pressure within the inlet, hence the notion “aerothermopressor”.
    The use of spray injection for inlet pre-cooling was the basis for the paper “The Transition Engine: A Combined-Cycle Engine Concept for SSTO/Trans-atmospheric Vehicle Applications” (AIAA 95-2480), which I wrote and presented at the 1995 AIAA Joint Propulsion Conference. Although in principle, with a sufficiently high ratio of injectant flowrate-to-air flowrate any particular stagnation temperature can be maintained, at some point the injectant flow becomes the dominant flow through the inlet. In my paper, it seemed that around Mach 5 represented the upper limit where spray injection made sense. I should also point out that in my paper, the spray injection was used for maintaining a particular inlet temperature, and not for trying to take advantage of the “aerothermopressor” effect.
    The question that I find interesting is not “why are rockets a preferred acceleration engine?”, but “where would airbreathing engines EVER be a preferred acceleration engine?” Answering that question has the potential to be fruitful, even if the answer is “never”.
    However, I don’t’ think it is. I claim, not without basis, that airbreathing engines, and in particular the ATR, may have a role as an airbreathing component to suborbital, and possibly orbital launch vehicles. This of course, cannot be a blanket statement. Such a statement is loaded with all sorts of assumptions and caveats.
    As I stated in my own post, I don’t necessarily believe that airbreathers have been oversold, I just think that we have come to the conclusion that the effort that is required to get them to work isn’t worth the benefits they provide.
    The honest answer to the question of the role of airbreathers for space launch vehicles, is that we really don’t know whether there’s a role for them or not. But to say that they have no role is, in my opinion, shortsighted. One might end up missing out on a useful approach. If I didn’t think they didn’t offer some possible advantages, I wouldn’t be working on them, but I think it’s fair to say that I am in the minority.
    I do feel that there is an upper limit in terms of flight speed, beyond which airbreathing propulsion simply does not make any sense. In my opinion, that limit is probably around mach 5 or 6. Beyond that flight speed, the technical challenges become very difficult, and the benefits are rapidly fading away. Dr. John Whitehead of Lawrence Livermore National Laboratory (and of rocket piston-pump fame) provided an very interesting analysis on the utility of airbreathing engines in his 2007 AIAA JPC paper: “Airbreathing Acceleration Toward Earth Orbit”, AIAA 2007-5837. He concluded that the maximum flight for airbreathing propulsion is limited by the ratio of air capture area to vehicle drag area, and that for equal areas, this maximum flight speed is about Mach 6. Dr. Whitehead’s background tends more towards rocket propulsion and he is by no means an airbreathing enthusiast, so I felt he provided a reasonable, objective assessment of airbreathing propulsion for launch vehicles.
    So what are we left with? Is there an airbreathing propulsion option that can take us from zero to mach 5 or 6? Does it have a sufficiently high T/W and Isp to justify its use? Can such a system be cost effective, and does it provide operational margin? Maybe there is such an engine, and more importantly, maybe there’s a vehicle configuration that can take advantage of these attributes. We’ll see.

  16. Check this turboflow option (liquids & gas):
    The Imploturbocompressor.
    And I have other project, the Gearturbine, details at:

    Tip Info / New Technology Submission – Gearturbine – Atypical

    http://gearturbine.260mb.com/

    YouTube Video; Atypical New * GEARTURBINE / Retrodynamic

    GEARTURBINE -Atypical Combustion Turbine Engine, -State of the Art, -New Thermodynamic Technology, -With Retrodynamic “Dextrogiro vs Levogiro” Effect, is when the inflow direction moves is against of the circular rotary dynamic, RPM Rotor Move VS Inflow Conduits Way, making in a simple way a very strong concept of power thrust, a unique technical cuality. -Non Waste, parasitic losses form-function engine system for; cooling, lubrication & combustion; -Lubrication & Combustion inside a conduit radial position, out way direction, activated by centrifugal force (centrpetal to in), -Cooling in & out; In by Thermomix flow & Out by air Thermo transference, activated by the dynamic rotary move, -Increase the first compresion by going of reduction of one big circunference fan blades going to, -2two very long distance cautive compression inflow propulsion conduits (like a digestive system) (long interaction) in perfect equilibrium well balanced start were end like a snake bite his own tale, -Inside active rotor with 4 pairs of retrodynamic turbos (complete regeneration power system), -Mechanical direct “Planetary Gear” power thrust like a Ying Yang (very strong torque) (friendly loose friction) 2two small gears in polar position inside a bigger shell gear, wide out the rotor circunference were have much more lever power thrust, lower RPM in a simple way solution for turbines, to make posible for a some new work aplication (land). -3 Stages of inflow turbo compression before the combustion. -3 points united of power thrust; 1- Rocket Flames, 2-Planetary Gear & 3-Exhaust Propulson, all in one system. -Combustion 2two continue circular moving inside rocket Flames, like two dragons trying to bite the tail of the opposite other. -Hybrid flow system diferent kind of aerolasticity thermoplastic inflow propulsion types; single, action & reaction turbines applied in one same system, -Military benefits, No blade erosion by sand & very low heat target profile. -Power thrust by barr (tube); air sea land & generation aplication, -With the unique retrodynamic technical cuality of “dextrogiro vs levogiro” effect is when the inside flow moves against of the rotor moves making a very strong concept, RPM Rotor Move VS Inflow Conduits Way (an a example is like to move the head to the side of the strike ponch) -A pretender of very high % efficient power plant looking to make posible a cheap electrolysis. -Patent; Dic 1991 IMPI Mexico #197187

  17. john hare says:

    Carlos,

    I just glanced at the first page or so in the link and don’t have time to fully check it out now. I’ll respond in a few days when I get back in town.

  18. SimonDM says:

    Supersonic impulse turbines is what you’re thinking of. They have been thought of and some designed, but as far as I understand efficiency is typically low because of flow separation issues.
    Water particles and turbomachinery is also not a good mix. Putting the turbine up front and the compressor behind is also nothing new, see United States Patent 4224790 which is massively interesting to me. You need however lots more cooling than you can achieve with water to make that work, such as liquid hydrogen. The problem is that such an engine produces very little thrust at zero speed.

    A solution to that problem would be to let the heated hydrogen drive a tip turbine located around the compressor, producing useful work and increasing compression of the air, and to have an additional combustion chamber in front of the turbine only to be used at low speeds. In this way, you would have only one single heat exchanger contrary to the multitude envisioned by ReactionEngines or the two envisioned in the ATREX engine, and you’d not have a heat exchanger in the main combustion chamber as in the atrex engine. Existing materials might do for heat exchangers and turbomachinery. Plus the efficiency would be higher than both the SABRE and ATREX engine on a first thermodynamic glance.

  19. john hare says:

    It’s not a classic supersonic impulse turbine as found on gas generator rockets. It is an intake turbine using RPM to produce variable ramp angles scheduled with airspeed.

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