Affordable Triprop Pumped Engine

guest blogger john hare

It is interesting how many people tend to think of current rocket engines as being mature or nearly so. In this ongoing series of thought experiments and looney tune ideas, I suggest that current engines are still in the infancy of development. A Kero/LOX engine with a T/W of 1,000 and sea level Isp of 350 is theoretically possible. Just jack the pressure to a kilobar or so without frying the buckytube construction is all. Piece of cake.

Even without invoking unobtanium, it is possible to see a few areas of possible improvement. The pumped monoprop I suggested a few weeks ago has part of a possible solution. Put all the high pressure components in a single pressure vessel without a lot of plumbing. This is another page in that direction for a triprop engine. Some of the comments grazed my intent for a high performance engine, with a couple hitting parts of the real target.

Hutzel and Huang has a very interesting breakdown of component mass in chapter 2. Thrust chamber mass down with increasing operating pressure, while turbopump and plumbing mass rises. They project a 750k thrust chamber at about 3,000 pounds mass at 3,000 psi. It drops to just over 2,000 pounds at 7,000 psi, but climbs to over 3,500 pounds at 2,000 psi. Pump and plumbing mass increase in a straight line with increasing pressure to cancel the gains from higher pressure. Optimization for the cost is no object engines is somewhere in the 3,000 psi range. Everyone knows that Isp increases with increasing pressure, it’s just that the increased pressure comes with a much increased price tag.

So if the high pressure plumbing can be eliminated, and a turbopump cycle can be found that does not increase mass quickly with higher pressure, it seems possible to get an affordable engine operating at a T/W of 150+, and a sea level Isp of 300+. While it is obviously theoretically possible technically, it may not be financially possible using conventional design practice.

Unconventional design practice comes with the price tag of uncertainty added to the dollar cost. Uncertainty is a variable that is unacceptable much of the time, like when the product has to be delivered or the company folds. Low performance sometimes comes with an even higher price tag. When a relatively low performance architecture is chosen just because it is low technical risk, sometimes the engineering effort and margin shaving required to make it perform the job are more expensive than if more technical risk is accepted up front. Seeking a middle ground between doing it just because it is new, and doing it just because it is proven seems desirable.

I propose taking the same spherical housing technique that I suggested for the pumped monoprop and use it nearly unchanged for a triprop. By using a small quantity of hydrogen to keep the turbine blades cooled, a very hot and high pressure Kero/LOX preburner can be used. The hydrogen is pumped through the turbine blades by the turbine blades even as it cools them. With the turbine tip moving around three times the tip speed of the impellers, the hydrogen is pumped to a head height nine times that of the impellers, which cancels (for pumping purposes) the low density of the hydrogen compared to the denser kerosene and oxygen.

The turbine tips at 2,000 feet per second drive an impeller at 700 fps for the kerosene and 600 fps for the oxygen. With the kerosene pressure drop through the cooling jacket and injectors, oxygen pressure drop through the injectors, and hydrogen heating in the turbine blades lowering its’ density, net liquid pressure averages a bit under 2,000 psi in the upper preburner section for all three propellant components. With hydrogen cooling of the turbine blades, the preburner can run much hotter than normal practice which provides much more power to drive the turbine. With more available power to drive the turbine, a lower pressure drop than normal is required to drive the system. Realized lower chamber rocket section should be in the 1,400-1,500 psi range. The oxygen that was not used in the preburner is injected into the turbine exhaust to bring the temperatures up to full rocket performance. Sea level Isp should be in the 300 range.

Some Russian kerosene engines run at higher pressures and have higher Isp without needing hydrogen for blade cooling. I think it is possible that a higher T/W can be achieved  with an engine series even cheaper than their offerings. In vacuum, the higher pressure they offer will affect the Isp a bit less.

I recycled the monoprop sketch to show the family history, or because I was lazy, pick one. As before, the impellers  use bowl volutes for simplicity and cost. The LOX feeds directly into the preburner injector and sends some percentage to the main engine injectors. The kerosene cools the turbine nozzles and bearings before entering the impeller. After the impeller, it enters the  two pass regenerative cooling passages before routing back to the preburner injector housing. The hydrogen in blue enters the turbine through the hub and cools the turbine disk and blades before ejection through the turbine tips and cooling holes in the turbine surface. It provides some film cooling of the main chamber in the injection zone.

The turbine disks and impellers I have handled were a fraction of the mass of the housings and plumbing associated with them. By eliminating many of the components like gearboxes and asymetrical housings usually required by a turbopump system, and consolidating functions with the actual engine the turbopump system serves, a 10k engine could reduce to 60-70 pounds without requiring massive design and development effort. I don’t think it is likely to get much better capability than suggested here, while retaining a conventional layout. More on unconventional later.

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johnhare

johnhare

I do construction for a living and aerospace as an occasional hobby. I am an inventor and a bit of an entrepreneur. I've been self employed since the 1980s and working in concrete since the 1970s. When I grow up, I want to work with rockets and spacecraft. I did a stupid rocket trick a few decades back and decided not to try another hot fire without adult supervision. Haven't located much of that as we are all big kids when working with our passions.
johnhare

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johnhare

About johnhare

I do construction for a living and aerospace as an occasional hobby. I am an inventor and a bit of an entrepreneur. I've been self employed since the 1980s and working in concrete since the 1970s. When I grow up, I want to work with rockets and spacecraft. I did a stupid rocket trick a few decades back and decided not to try another hot fire without adult supervision. Haven't located much of that as we are all big kids when working with our passions.
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16 Responses to Affordable Triprop Pumped Engine

  1. Tim says:

    Does Hurtzel and Huang count the cooling passages as part of the thrust chamber or part of the plumbing? Given that its “outside” the pressure chamber and at high (?) pressure, is there something about cooling passages that mean they don’t cause as great a mass problem as separate plumbing (I figure a mass hit comes from the requirement for thicker thrust chamber walls to accomodate the passages)?

    As I understand it, triprops are supposed to run heavy on kero and light on LH2 at low altitude, increasing the LH2 content and decreasing the kero as altitude increases. Is that going to cause cooling problems ie. A hot turbine at low altitude and a hot bearing at high altitude? Or at least lead to difficult compromises?

    Is the propellant coming out of the turbine/main injector in a ring shape? If it is, it might make a nice annular injector for an annular aerospike, if you can figure out how to cool the thing.

  2. john hare says:

    Tim,

    Does Hurtzel and Huang count the cooling passages as part of the thrust chamber or part of the plumbing? Given that its “outside” the pressure chamber and at high (?) pressure, is there something about cooling passages that mean they don’t cause as great a mass problem as separate plumbing (I figure a mass hit comes from the requirement for thicker thrust chamber walls to accomodate the passages)?

    Cooling passages are part of the chamber. They are actually inside the walls of the chamber except that I sketched it wrong. From the hot side there would be a thin layer of conductive metal, then cooling passages with ribs, and finally a thick structural shell on the outside. I was just trying to convey the fluid directions without drawing all the passages in the proper place like I should have.

    As I understand it, triprops are supposed to run heavy on kero and light on LH2 at low altitude, increasing the LH2 content and decreasing the kero as altitude increases. Is that going to cause cooling problems ie. A hot turbine at low altitude and a hot bearing at high altitude? Or at least lead to difficult compromises?

    Triprops work best as you describe. I was suggesting that only enough LH2 be used to cool the turbine for the whole flight profile. 1% LH2 or less at any given time. This is barely a triprop as it is laid out, with no mixture shift and the LH2 unreacted with LOX. It will provide good reaction mass though.

    Is the propellant coming out of the turbine/main injector in a ring shape? If it is, it might make a nice annular injector for an annular aerospike, if you can figure out how to cool the thing.

    It does come out in a ring. I think it needs a little more volume for combustion than an annular unit would provide. That would be an interesting design problem, and a good argument for a larger diameter turbine for lower rpm operation. I’ll put that on the list.

  3. jsuros says:

    If tip speed drives the design, and therefore larger sized engines have lower rpm, how far up can you scale this engine? Is there a sweet spot somewhere?

    Also, speaking as a Black Horse fan, could you use H2O2 and JP-5 as the propellants for this engine?

    One last question. In your original article, you based this design around low inlet pressures. How do the numbers scale up if you were to feed in propellants at high pressure?

  4. john hare says:

    Jsuros,

    If tip speed drives the design, and therefore larger sized engines have lower rpm, how far up can you scale this engine? Is there a sweet spot somewhere?

    Combustion instability would be my major worry at larger sizes considering all the problems they had getting the F1 stable. If that does not stop growth, then I see no technical reason not to grow it as large as the mission requires, including a single engine Saturn 5 class vehicle. Finding the sweet spot would require analysis beyond my capabilities. I am looking for generic system layouts that have the possibility of economical development more than a specific size.

    Also, speaking as a Black Horse fan, could you use H2O2 and JP-5 as the propellants for this engine?

    Assuming the cycle works as diagrammed, that would be a good choice as the decomposition could be completed in the precombustion section prior to main chamber.

    One last question. In your original article, you based this design around low inlet pressures. How do the numbers scale up if you were to feed in propellants at high pressure?

    It increases the pumping reliability with cavitation elimination, and increases the total available pressure operation. For Isp and engine T/W it is a major win. The question is how much mass do you pay to get that higher pressure inlet pressure. If you are simply raising the initial operating pressure of the same tanks, it is probably a clear win. If you are massively increasing the tank mass to get the pressures, probably not. If you are running a gas generator type inducer, maybe.

  5. Gary C Hudson says:

    Combustion stability is far less of an issue with a spherical chamber.

  6. jsuros says:

    Following up on the “higher pressure input leads to better performance” riff…

    Flowmetrics claims their pistonless pump can pull propellants from a low pressure tank and deliver them to a combustion chamber at 1000 psi with a T/W of 300-1000 depending on design assumptions.

    Suppose you used one of these pumps as a precompressor to your triprop design. Can you add the 1000 psi to the 1400 psi chamber pressure in your base design to get 2400 psi, or is it more complicated?

  7. James says:

    Hi John,

    If what you are looking for is high ISP from a hydrocarbon fuel, check out these links:

    http://www.wipo.int/pctdb/en/wo.jsp?IA=US1999017720&DISPLAY=DESC

    http://www.wipo.int/pctdb/images4/PCT-PAGES/2000/312000/00009880/00009880.pdf

    This is a patent from 1999 describing a ROMP based fuel (poly-dicyclopentadiene) combined with lithium aluminum hydride. In particular, check out the ISPs in the second link. For LOX as the oxidizer, the ISP is around 450, for peroxide, it is around 430. As you probably know, these kinds of ISPs are typically only achieved with LH. Since ROMP fuels are solid, this would be a hybrid. I’ve also seen some work from the 60’s on combining metal hydrides with kerosine that showed similar results (i.e. substantially improved ISP over the base fuel), but with enough metal hydride, the kerosine becomes thixotropic and the fuel is a gel, which has its own problems in terms of pumping.

    jak

  8. john hare says:

    Jsuros,

    Following up on the “higher pressure input leads to better performance” riff…

    Flowmetrics claims their pistonless pump can pull propellants from a low pressure tank and deliver them to a combustion chamber at 1000 psi with a T/W of 300-1000 depending on design assumptions.

    If Steve and the gang can deliver performance at that level, the trades could easily lead to elimination of the turbo system we are discussing. At Space Access, they use them to pump margaritas.

    Suppose you used one of these pumps as a precompressor to your triprop design. Can you add the 1000 psi to the 1400 psi chamber pressure in your base design to get 2400 psi, or is it more complicated?

    Yes you can do that and get serious performance improvements. By using a Flowmetrics unit as an inducer, you go from cavitation being a major problem and limit, to material strength of the impellers being the limit. 6,000 or more psi at the pump becomes possible with net in the chamber of ~4,000 psi.

    XCOR has a composite material they developed for LOX tanks. If it can be worked for impellers, impeller speed limits increase to ridiculous levels, with theoretical pressure into five digits.

  9. johnhare johnhare says:

    James,

    What I look for is cheap ways of getting expensive performance. If a relatively simple layout can let newspace match the performance of the best current engines. So my first question is, what do these propellants cost? In price per pound of propellant, indevelopment cost of hardware if higher, and safety. Regulatory problems if any.

    I got exited about the boron based stuff a few years ago until I found out how dangerous it was. I couldn’t get the second link to work and don’t really know anything useful about the fuel you recomended. Metals do have some useful properties in fuel, especially aluminum.

  10. I know this is a discussion about tripop engines, so my comment/question is tangential, but here goes:

    Would injecting a hydrocarbon propellant into the exhaust flow of a LOX/H2 engine (essentially an “afterburner” effect) provide some of the benefits attributed to triprop propulsion? I understand that a typical LOX/H2 mixture ratio is 6-to-1 LOX to H2 (does that sound right?), so I’m assuming an oxygen-rich exhaust (combustion is never complete). I would imagine the hydrocarbon fuel injectors would be in the engine throat.

    Another, similar option would be for the launch vehicle to have an “afterburner skirt” that would be a lower expansion ratio nozzle in and of itself, and would extend beyond the nozzle(s) of the engine(s). RP-1 or another hydrocarbon propellant would be injected at this point into the engine exhaust in the “skirt.” The skirt/sleeve nozzle attachment could also have air inlets at the attachment point at the lower end of the fuselage, so that it would also be a primitive air-ducted rocket.

    You’d “lose the skirt” (pardon the expression) at the end of the boost phase, along with an external tank containing the propellant for the boost phase (they can both be recoverable). The launch vehicle is now all LOX/H2, and without the skirt, has a higher expansion ratio, AND it’s size and mass are reduced with the ejection of skirt and boost phase tankage.

    OR,

    Maybe I have no idea what I’m talking about and should just have another beer.

  11. john hare says:

    Roderick,

    I know this is a discussion about tripop engines, so my comment/question is tangential, but here goes:

    This is a discussion of technical means for reducing the cost of spaceflight. Anything that helps that goal is good.

    Would injecting a hydrocarbon propellant into the exhaust flow of a LOX/H2 engine (essentially an “afterburner” effect) provide some of the benefits attributed to triprop propulsion? I understand that a typical LOX/H2 mixture ratio is 6-to-1 LOX to H2 (does that sound right?), so I’m assuming an oxygen-rich exhaust (combustion is never complete). I would imagine the hydrocarbon fuel injectors would be in the engine throat.

    The 6-to-1 mixture ratio is a bit rich with some unburned hydrogen in the exhaust. Stoichiometric would be 8-to-1. To get the effect you are suggesting, mixture would need to be 9-to-1 or better. Hydrogen engines actually get better Isp with a richer mix yet because it drops the average molecular weight of the exhaust. If you were to inject oxygen into the exhaust, it would burn a little hotter and kill Isp with both the increased average molecular weight and the unburned oxygen that adds no energy and has much higher molecular weight.

    Another, similar option would be for the launch vehicle to have an “afterburner skirt” that would be a lower expansion ratio nozzle in and of itself, and would extend beyond the nozzle(s) of the engine(s). RP-1 or another hydrocarbon propellant would be injected at this point into the engine exhaust in the “skirt.” The skirt/sleeve nozzle attachment could also have air inlets at the attachment point at the lower end of the fuselage, so that it would also be a primitive air-ducted rocket.

    I’m having trouble visualizing some of your idea here. The air inlets make some sense, the afterburner skirt I can’t figure. Did you check out Jons’ post on augmented nozzles a while back?

    You’d “lose the skirt” (pardon the expression) at the end of the boost phase, along with an external tank containing the propellant for the boost phase (they can both be recoverable). The launch vehicle is now all LOX/H2, and without the skirt, has a higher expansion ratio, AND it’s size and mass are reduced with the ejection of skirt and boost phase tankage.

    It seems to me that you want to keep the skirt for maximum expansion ratio, so I must be missing your intent.

    OR,

    Maybe I have no idea what I’m talking about and should just have another beer.

    This year has increased my beer consumption by about 50%. Normally it’s around a case a year.

  12. Thanks for the reply. The whole point is whether or not an “afterburner’-like concept would work for a simple tri=propellant system.

    It would be useful for me to illustrate it, and, when I can find some time, I can do that. I should check Jon’s augmented nozzle post. How far back was that?

    Also, thanks for the technical feedback on the “fuel-injection” idea. I had suspected that the LOX ratio would have to be higher, but wasn’t sure. I also see that the idea may not work.

  13. john hare says:

    Roderick,

    November 2007 for the thrust augmented nozzle post. It’s a good read. It is kinda afterburner like.

  14. John,

    I assume the 11/2007 post you refer to is this one:
    http://selenianboondocks.com/2007/11/thrust-augmented-nozzles/
    Thanks

    I also liked the one about the NK-33-34 engines:
    http://selenianboondocks.com/2007/11/random-thought-thrust-augmented-aj26-60/

    I had a number of interesting discussions with Aerojet guys back in the 80’s. Guys like Dick Morrison and Rudi Beichel. Cool stuff.

  15. Oh, also:
    re: http://selenianboondocks.com/2007/11/thrust-augmented-nozzles/

    That was exactly what I was getting at with my notion of an “afterburner.” I had never heard of this research. It makes good instinctive sense to me.

    As the post stated, it seemed, in theory, to be a great way to get an extra boost off the pad, plus the benefits of higher isp at altitude. This is what I was trying to get at, but fumbled.

    Thanks again for the links.

  16. Jonathan Goff Jonathan Goff says:

    Roderick,
    Yeah, the TAN stuff is definitely interesting. One important thing to remember about Aerospace is that it’s pretty common for good ideas to be independently thought up many times before they actually are reduced to practice. I know of at least one or two other groups who suggested similar ideas even further back than Aerojet’s TAN work. But Aerojet was the first one to actually build and test fire the concept. It’ll be interesting to see who actually flies the first TAN engine.

    ~Jon

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