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From: (Henry Spencer)
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 |

From: Henry Spencer <>
Subject: Re: engine reliability (was Re: Venturestar turnaround...)
Date: Tue, 28 Jan 1997 18:47:41 GMT

In article <>,
Michael P. Walsh <> 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."|

From: Henry Spencer <>
Subject: Re: NASA Engine Development
Date: Sun, 20 Apr 1997 04:13:24 GMT

In article <>,
Jim Kingdon <> 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        |

From: Henry Spencer <>
Subject: Re: Black Horse (was: Re: HST as justification for STS)
Date: Wed, 23 Apr 1997 12:15:34 GMT

In article <>,
Jacob McGuire  <> 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        |

From: Henry Spencer <>
Subject: Re: RL10 (was Re: Shuttle II...)
Date: Sun, 9 Jun 1996 04:28:36 GMT

In article <4ol5q0$ill$> 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 |

From: Henry Spencer <>
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$> (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 |

From: Bruce Dunn <>
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)
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

From: (Henry Spencer)
Subject: Re: RL-50
Date: Wed, 30 Jun 1999 17:56:25 GMT

In article <>,
Michael P. Walsh <> 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
weren't.                            |      (aka

From: (GCHudson)
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 <>
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.

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.
Reliable, low-cost transportation to low Earth orbit and beyond

From: (Henry Spencer)
Subject: RL10 (was Re: Aerial Propellant Transfer Revisited)
Date: Wed, 5 Jul 2000 13:44:51 GMT

In article <>,
Kirk Voelcker  <> 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

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
It should be shut down.  -- Phil Agre   |      (aka


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