From: henry@zoo.toronto.edu (Henry Spencer) Newsgroups: sci.space.tech Subject: Re: Soviet Reusable(?) engine Date: Fri, 27 Oct 1995 16:10:00 GMT Organization: U of Toronto Zoology Lines: 24 >>> Just to say for the umpteenth time what's been said before, Henry, if you >>> put together all of the flights of the DC-X in one, it has not >>> accumulated enough engine-on time to get to orbit, or anywhere near >>> orbit, or even, I don't think, particularly far out of the atmosphere. Just to say for the umpteenth time what I've consistently said in response: the off-the-shelf RL10 is rated for 4000s continuous firing without any maintenance (assuming you can somehow keep it fed with fuel). I don't remember how many seconds the high-time RL10 has on it, but it's a lot more than that. There is no reasonable doubt that once started, an RL10 can keep running almost indefinitely. What looked more worrisome was the combination of repeated starts, minimal maintenance, and storage in semi-controlled conditions; DC-X has fairly decisively settled that. It flew more often (averaged over its flight career, even including the long hiatuses) than any of the shuttle orbiters, and with orders of magnitude less babying. To reiterate my original point: almost any regeneratively-cooled liquid rocket engine is reusable, even if it was built for expendable vehicles. The design requirement for the F-1 was 20 starts and 2250s of firing; as part of the test program, six of them accumulated over 5000s each. -- The problem is, every time something goes wrong, | Henry Spencer the paperwork is found in order... -Walker on NASA | henry@zoo.toronto.edu Newsgroups: sci.space.policy,sci.space.history From: Henry Spencer <henry@zoo.toronto.edu> Subject: Re: engine reliability (was Re: Venturestar turnaround...) Date: Tue, 28 Jan 1997 18:47:41 GMT In article <32ED205C.75E1@pacbell.net>, Michael P. Walsh <mp_walsh@pacbell.net> wrote: >> ...Perhaps a couple of hundred RL10s have flown, >> and until fairly recently, the RL10 had never had an in-flight failure... >--- >I know there were failures of the Centaur propulsion system to reignite >and perform properly, resulting in the loss of the payload. I don't >know whether this was a failure of the RL-10 or the other parts of >the propulsion system or perhaps the control system. There were diverse problems with Centaur, but when the RL10s were given a chance, they worked. Atlas-Centaur 1 disintegrated during Atlas ascent due to Centaur structural problems caused by thermal contraction. AC-2 flew perfectly. AC-3 lost attitude control due to a failure in a shaft coupling of a hydraulic pump. The engines stopped when spin built up to the point that their LOX feed was cut off. AC-4, the first attempt at a two-burn Centaur flight, lost attitude control during coast and tumbled, but to the amazement of all the RL10s managed a successful restart (albeit with severely reduced performance) despite having their LH2 lines full of hydrogen gas instead of liquid. AC-5's Atlas lost thrust at an altitude of 1 meter, fell back and exploded. AC-6, a one-burn flight, was a complete success. AC-8's peroxide system leaked in coast, and there wasn't enough left to spin up the boost pumps (part of the vehicle, not the engines) to supply proper pressure at the engine inlets for restart. One RL10 nevertheless managed to restart and reached full thrust, but the other didn't, and the resulting tumble cut off propellant feed to the successful engine. (The boost pumps were a perennial source of trouble, and were eventually eliminated by design changes.) AC-10 launched Surveyor 1, a complete success. AC-7 launched Surveyor 2, perfect launch but the spacecraft failed. AC-9 finally flew a completely successful two-burn flight. Anyway, the picture after that remained generally the same -- the stage gave occasional trouble but the engines didn't -- until 1991 and 1992, when a pair of Centaurs failed because one engine of each didn't start, due to the failure of its pumps to spin up. The cause is tentatively thought to have been air or nitrogen ice in the turbopump. Changes were made, mostly in procedures rather than hardware, and the problem has not recurred. There was one later Atlas-Centaur failure due to an Atlas problem (the Centaur tried to make up for it but ran out of fuel trying). -- "We don't care. We don't have to. You'll buy | Henry Spencer whatever we ship, so why bother? We're Microsoft."| henry@zoo.toronto.edu Newsgroups: sci.space.policy,sci.space.history From: Henry Spencer <henry@zoo.toronto.edu> Subject: Re: NASA Engine Development Date: Sun, 20 Apr 1997 04:13:24 GMT In article <m2912l3bh2.fsf@harvey.cyclic.com>, Jim Kingdon <kingdon@harvey.cyclic.com> wrote: >> The question that arises is in just what manner does Lewis >> take credit for the RL-10? > >Here is a quote. If someone has some historical perspective to offer >(like what this tells us about the relationship of government to >industry, or whatever), go ahead... While Lewis undoubtedly would like to take credit for the RL10 :-), and they were later involved in the management of Centaur, their contributions to the RL10 seem to have been fairly indirect. Probably their biggest role in it was simply convincing the NASA brass that LOX/LH2 was feasible as a high-performance fuel combination. Although Pratt&Whitney certainly talked to the Lewis people with some frequency, their design seems to have been largely independent. Much of the RL10 pump hardware, a lot of the test facilities (including the LH2 plants), and considerable experience in liquid-hydrogen cooling came from their work on the engine for the then-highly-classified Suntan project (a hydrogen-fueled spyplane, which died young because of problems achieving adequate range). The RL10's expander cycle (absorbing engine heat in the hydrogen, and tapping the resulting warm hydrogen gas to run the pump turbine before injecting it) came directly from the Suntan engine. The methods used to design the nozzle were cited as a major P&W innovation. And the P&W injector -- the trickiest part of a rocket engine -- looked absolutely nothing like the Lewis injectors, with concentric injector elements (where Lewis used more conventional showerhead and impinging designs) and a porous-mesh faceplate (where Lewis used a conventional solid one). -- Committees do harm merely by existing. | Henry Spencer -- Freeman Dyson | henry@zoo.toronto.edu Newsgroups: sci.space.history,sci.space.policy From: Henry Spencer <henry@zoo.toronto.edu> Subject: Re: Black Horse (was: Re: HST as justification for STS) Date: Wed, 23 Apr 1997 12:15:34 GMT In article <QnKyJcy00iWS4Ds0cf@andrew.cmu.edu>, Jacob McGuire <mcguire+@andrew.cmu.edu> wrote: >>I believe you're thinking of a Russian engine, rather than the RL-10. >>Redesigning a LOX/hydrogen engine to use H2O2/kerosene would be a much >>bigger deal than redesigning a LOX/kerosene engine. > > "CH4/O2 offers the second highest Isp (385 s)... A significant >disadvantage of CH4/O2 is that no flight-rated engine is >currently available; development of one based on Pratt and Whitney RL-10 >engine technology would probably take three years and cost >on the order of $30 million." One reason to be reasonably confident in this low number is that the RL10 has already been run on LOX/CH4, with only minor modifications. That was done only experimentally, and you'd need a bit more work to certify that configuration for production use, but it shouldn't be hard. The RL10 has actually been run on a very wide variety of things. For fuels, hydrogen, methane, and propane have all been used; for oxidizers, they've fired it on LOX, liquid fluorine, and FLOX (LOX/fluorine mixture). Mixture ratio has been varied extremely widely, including successful tests running severely oxidizer-rich. And it's been throttled down to 1%. An amazingly tolerant engine. -- Committees do harm merely by existing. | Henry Spencer -- Freeman Dyson | henry@zoo.toronto.edu Newsgroups: sci.space.policy From: Henry Spencer <henry@zoo.toronto.edu> Subject: Re: RL10 (was Re: Shuttle II...) Date: Sun, 9 Jun 1996 04:28:36 GMT In article <4ol5q0$ill$2@mhade.production.compuserve.com> Michael P. Walsh <71544.2040@CompuServe.COM> writes: >>...But I have seen at least one >>supposedly orbit-capable SSTO demonstrator design powered by RL10s. > >The problem I have with an RL-10's clustered as an SSTO main >engine package is that operating the engine at sea level with a >conventional nozzle requires reducing the expansion ratio to >avoid flow separation. Indeed so, as witness the stubby little nozzle on the RL10A-5 variant built for DC-X. You probably need to do *something* about altitude compensation if you want to power an SSTO with RL10s, although it might be done in the vehicle itself rather than in the engines -- e.g., an aerospike using a ring of RL10A-5s as its chambers. >Using a pure expander cycle (all pump >energy received from heat transfer to the propellant) means >limited chamber pressure... Yup, this is why the Europeans rejected the expander cycle for the Ariane 5 engine -- they wanted to use a 100atm chamber pressure, and concluded that this was just too high for an expander cycle. >I believe the RL-10 is about 300 psi >(corrections accepted, and the chamber pressure is the key to the >whole discussion)... The original RL10A-1 was 300psi, but that was changed quite early to 400psi (via a throat restriction) for improved performance. >Getting Isp performance back up could be done >with a variable nozzle extension, at added weight and complexity... Note that a telescoping nozzle extension is already an option for the RL10 variant flying on the current Atlas, although it deploys fully before ignition. I'm told that P&W has tested an RL10 nozzle which telescopes during firing, and it worked, but as far as I know this was just an experiment and it's not an option for production engines at the moment. -- If we feared danger, mankind would never | Henry Spencer go to space. --Ellison S. Onizuka | henry@zoo.toronto.edu From: Henry Spencer <henry@zoo.toronto.edu> Newsgroups: sci.space.tech Subject: Re: RLV engines (was Re: X-33 Concepts: Lockheed, Mac Dac, Rockwell) Date: Mon, 24 Jun 1996 05:00:28 GMT In article <4qboqu$t1r@ns.hcsc.com> andyh@hcxio.hdw.hcsc.com (Andy Haber) writes: >>...Most ELV engines are, despite their application, reusable, because >>they have to be developed and tested. The F-1 was specified for 20 starts >>and 2250s of life, the J-2 for 30 and 3750s. Six F-1s ran over 5000s each >>as part of the service-life tests. DC-X's RL10s looked "pristine" after >>20 starts; the RL10 is nominally rated for 10 starts and 4000s of firing. > >Still even these numbers are not all that good... True, but that's partly because nobody has thought it worth testing these engines for still longer lives. (Mind you, I don't have numbers for the RL10 "fleet leader" test engine -- I wouldn't be surprised if it has rather more than 4000s on it.) Some of them -- my money would be on the RL10 in particular :-) -- very probably have much longer lives than their current test history is adequate to demonstrate. >>Unfortunately, it probably can't go far enough. Rocketdyne's own estimate >>was that, with a *lot* of work, you could probably get SSME maintenance >>costs down to $750k/engine/flight, which is unsatisfactory if you're aiming >>for really large cost reductions. > >This seems like somewhat of a "disconnect" to me. If a RL-10A-3-3A will >do 10 starts and 4000 seconds with no maintenance, then why, after a *lot* >of work, does a SSME need 3/4 of million dollars worth of maintenace after >1 start and 500 seconds of use? ... I don't know what the breakdown of the $750k is; my info on this is secondhand. However, the RL10 is a much simpler engine, using much more ordinary materials, running at much lower pressures and temperatures, and quite substantially overbuilt in certain areas. It has major inherent advantages when durability and maintenance requirements are the issue. -- If we feared danger, mankind would never | Henry Spencer go to space. --Ellison S. Onizuka | henry@zoo.toronto.edu From: Bruce Dunn <bruce_dunn@mindlink.bc.ca> Newsgroups: sci.space.tech Subject: Scaled RL-10 for SSTO? Date: Wed, 26 Jun 1996 22:22:17 -0700 I was asked by mail: >But why doesn't the RL10 need to be designed to operate in >an atmosphere? It needs to take off from the surface just >as much as the shuttle does. Also the external pressure >difference between being at sea level and space is only >15lbs/sq in so why did they have to go to such a huge >chamber pressure? Obviously there were other engineering >reasons involved. (ie why go from the RL10's 30 atms to >200 atm. just to correct for 15 lbs of air pressure). The reason of course that the RL-10 doesn't need to be designed to operate in an atmosphere is that it was purpose built as an upper stage engine. As a non-handwaving response to the rest of the question, I have done some rough calculations with a specific impulse program and some nozzle equations to show the difficulties in simply scaling up an RL-10 for use in a SSTO. These are basically "back of the envelope" calculations, but are instructive. Consider trying to make an equal thrust replacement for an SSME by scaling up an RL-10. SSME RL-10 Scaled RL-10 Mass, kg 3177 138 ???? Length, m 4.24 1.78 9.50 Throat Diameter, m 0.27 0.131 0.69 Exit Diameter, m 2.39 1.02 5.44 Expansion ratio 77.5 61 61 Chamber pressure, atm 204 32.2 32.2 Exit Pressure, atm 0.200 0.040 0.040 Vac Thrust kN 2091 73.4 2091 SL Thrust, kN 1668 xxx xxx Vac Isp 454 444 444 SL Isp 368 xxx xxx The SSME figures describe a complex, compact, high pressure engine designed both for operation in the atmosphere, and for operation in a vacuum. The RL-10 figures are for the model RL-10A-3-3A. This is a relatively simple low pressure engine designed originally for use in the Centaur upper stage. The "scaled RL-10" is an RL-10 stretched to give the same vacuum thrust as an SSME. It is a huge engine - since everything is operating at lower pressures than in an SSME, everything has to be correspondingly bigger to develop the same thrust. Even if it could be built and operated, it might be questionable whether the engine is compact enough to use in a multi-engine installation (consider the area swept out by gimbaling an engine nearly 10 meters long). The expansion ratio of the RL-10 or stretched RL-10 is nearly that of the SSME, and the delivered Isp in a vacuum is also nearly that of the SSME. However, the RL-10 engines cannot be operated at sea level as the nozzles are too big (indicated by "xxx" in the table). One of the key parameters to look at is the exit pressure (the pressure of the exhausted gas at the nozzle exit). On the SSME, the pressure is already markedly lower than one atmosphere. Most first stage engines are not operated with exit pressures below about half an atmosphere. A pressure lower than atmosphere gives Isp losses (the balance of pressures in the bottom part of the nozzle at sea level actually gives a net force in the opposite direction to that desired). The lower than atmospheric pressure also risks engine-destroying flow separation, in which the flow detaches unevenly from the walls of the nozzle and atmospheric pressure "sneaks into the gap". On an unmodified RL-10, using a 61 to 1 expansion ratio (by area) results in an approximately 800 to 1 expansion ratio by pressure (ratio of chamber to exit pressures). This puts the exit pressure at nearly zero. Even if the flow did not separate, the calculated thrust and Isp while operating in the atmosphere are also near zero. The "forward facing" thrust of the engine is nearly completely counterbalanced by the fact that there is a huge net backward force created by the pressure imbalance across the very large exit plane area (which is far bigger for a scaled RL-10 than for an SSME) Leaving aside variable geometry nozzles, the only thing that can be done to make the scaled RL-10 operable in the atmosphere is to increase the chamber pressure or cut back the nozzle (as was done for the engines of the DCX). There are limits however to how much can be done with an expander cycle engine in developing more pressure. Already in the RL-10, all the energy available from regenerative cooling is being applied to generate pumping power. If more power is needed, more heat energy is needed. This then starts to complicate the chamber, as engineers play tricks such as increasing the surface area of the chamber to give more heat flux. After exhausting the possibilities of the expander cycle engine, the only alternative to get even higher chamber pressures is to go to a different cycle (gas generator, staged combustion etc.). At this point, you no longer have a "simple, low stress, scaled RL-10", you have a European Vulcain, a Japanese LE-7, or an SSME. So how about cutting the nozzle back. If the scaled RL-10 is given a stubby 17 to 1 expansion ratio nozzle rather than the original 61 to 1 version, the exit pressure is the same as that in the SSME (and we know that the SSME can operate in the atmosphere). However, Isp suffers horribly. In space the SSME has an Isp of 454, and at sea level it is still 368. The 17 to 1 expansion version of the RL-10 has a vacuum Isp of 428 (down from 444 because the nozzle has been shortened). More critically, the sea level Isp is a miserable 255. This is because of the pressure differential losses across the still very large exit plane. The sea level performance of the engine can be considerably improved by cutting the nozzle back even more to make the exit plane smaller while at the same time boosting the exit pressure. A 10 to 1 nozzle boosts sea level performance up to an Isp of 329. However, at the same time it further degrades the vacuum performance to an Isp of 404. No matter how you play with a fixed nozzle, you can't at the same time have both good vacuum performance, and good sea level performance. The lower the engine pressure, further apart are the optimum expansion ratios for sea level and vacuum operation. But operation both at sea level and in a vacuum is just what an SSTO demands. If you want the operational simplicity and long life of the RL-10 low pressure expander cycle, you will just have to get busy and design some form of altitude compensating nozzle (telescoping nozzle sections, jet engine style petal exits, aerospikes, expansion-deflection engines, etc.). If you dig in your heels and demand a conventional fixed nozzle because that where the bulk of our experience is, then you are just going to have to live with some form of pumping system that generates several times the pumping power per unit mass as does the RL-10 powerhead. From: gchudson@aol.com (GCHudson) Newsgroups: sci.space.policy Subject: Re: RL-50 Date: 30 Jun 1999 03:50:38 GMT Tim Johnson wrote: >Good point. The RL10B-2 is up to almost 25,000 pounds thrust, >and its failure on Delta 3 may show that Pratt has already >crossed the RL10 limit! Not so. 35k versions of the RL10 have been proposed (and hardware tested) for many years. RL10s have run well above 25 on the stand (once indavertently). P&W have also propsed 50K versions in the past, but those have had little in common with the current RL10 hardware. The Delta failure is almost certainly a QA problem. Gary C. Hudson Newsgroups: sci.space.policy From: henry@spsystems.net (Henry Spencer) Subject: Re: RL-50 Date: Wed, 30 Jun 1999 17:56:25 GMT In article <3779638A.8DB80394@pacbell.net>, Michael P. Walsh <mp_walsh@pacbell.net> wrote: >To me, this seems a very tricky piece of engineering. The >RL-10 started out as a 15,000 lb. thrust engine and through >various upgrades is now up to about 20,000 lbs. I suspect >that is getting close to the limit... Actually, if I recall correctly, early in its development the RL10 was briefly meant to be something like a 30klb engine. Part of the reason why it has been so robust and well-behaved is that it is somewhat overbuilt for a 15klb engine. As Gary has already noted, it has been run at rather more than 20klb experimentally, without any particular difficulties. -- The good old days | Henry Spencer henry@spsystems.net weren't. | (aka henry@zoo.toronto.edu) From: gchudson@aol.com (GCHudson) Newsgroups: sci.space.policy Subject: Re: RL-50 Date: 01 Jul 1999 03:59:03 GMT Mike Walsh wrote: >You may have missed one thing I read that P&W expects to be able to produce the RL-50 engine for about the same cost as the RL-10. That is quite a goal for an engine that should have a larger combustion chamber and nozzle and higher capacity turbomachinery. I almost said larger turbomachinery, but more efficiency there might be one way they hold down costs. I have no inputs from P&W or anywhere else but Aviation Week and what I see on-line. Come on, knowledgable people, tell me where I am wrong.< The thrust rating of a pump fed engine is largely unrelated to its cost. The price of an SSME has varied by a factor of two, purely on minor manufacturing changes; the RL10 has done the same from over $4 million in 1986 to half that today. I've never found a correlation between thrust and size which was not tied to production rate and complexity. BTW, the LH2 pump of the RL10 was developed for the SUNTAN jet engine (see LH2 as a Propulsion Fuel, a NASA SP) and can flow more than twice the LH2 that the 15k RL10 requires. The turbine nozzles were blocked off to reduce pump flow from the jet engine version. Gary C. Hudson From: Bruce Dunn <bpdunn@home.com> Newsgroups: sci.space.policy Subject: Re: RL-50 Date: Thu, 01 Jul 1999 04:42:23 GMT For interest, the Pratt and Whitney web site indicates that the RL-10 engine currently comes in versions which have thrusts as high as 24,750 lbs. The site also has a nice cutaway picture of the engine, nested in its extendable nozzle. http://www.pratt-whitney.com/engines/gallery/rl10.html Henry Spencer wrote: > > Actually, if I recall correctly, early in its development the RL10 was > briefly meant to be something like a 30klb engine. -- Dr. Bruce Dunn General Astronautics Canada, Vancouver B.C. http://www.genastro.com/ Reliable, low-cost transportation to low Earth orbit and beyond Newsgroups: sci.space.tech From: henry@spsystems.net (Henry Spencer) Subject: RL10 (was Re: Aerial Propellant Transfer Revisited) Date: Wed, 5 Jul 2000 13:44:51 GMT In article <3962C32A.A425CF90@earthlink.net>, Kirk Voelcker <anaxagoras@earthlink.net> wrote: >> I>The RL-10 has been run on methane, yes? >> And propane. Only minor modifications were required... > >Do you have a URL or an ISBN that lists all the fuel combinations used by the >RL-10? Unfortunately, no... The closest thing is the RL10 paper in NASA CP-3112, "Space Transportation Propulsion Technology Symposium", 1990, which talks a bit about alternate propellants and growth versions. (NASA publications do not have ISBNs, sorry.) And even it doesn't mention some of the odder things you can find in obscure NASA reports, e.g. the brief tests with FLOX/1-butylene or the experiments at oxidizer-rich mixture ratios. Nor is it current enough to discuss the short-nozzle RL10 variant built for DC-X. The definitive RL10 history paper has yet to be written. -- Microsoft shouldn't be broken up. | Henry Spencer henry@spsystems.net It should be shut down. -- Phil Agre | (aka henry@zoo.toronto.edu) |
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