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From: Henry Spencer <>
Subject: Re: alternative SSTO rocket fuels
Date: Fri, 16 Feb 1996 05:53:45 GMT

In article <> Ruediger Klaehn <> writes:
>Why do they plan to use LH2/LO2 as a propellant for the X-33? ...

It's a combination of the limited choice of available engines, and the
mythology of liquid hydrogen (which was seen as a wonder fuel for so long
that few people have looked hard at its disadvantages). 

>...A better propellant would be LH2/LF2, since you have a
>higher specific energy and therefore a higher exhaust velocity, so you
>need a lower weight ratio...

Apart from the cost and safety problems of fluorine, and the toxic 
exhaust, there are no large LH2/LF2 engines in existence.

>...Another reaction with
>high specific energy is SiH4+2O2 -> SiO2+2H2O. The specific energy of
>this reaction is slightly higher than that of the H2/O2 reaction, but,
>more important, liquid SiH4 should have a significantly higher density...

Unfortunately, there is more to the problem of Isp than specific energy.
You also need to look at the efficiency with which the thermal energy is
converted to exhaust kinetic energy.  This is a strong function of the
exhaust composition, and is the reason why almost all propellant
combinations give better results when run significantly fuel-rich
(in the case of LH2-based combinations, very fuel-rich).  Any high-melting
component in the exhaust is a major disaster for energy conversion -- it
contributes little or nothing after it condenses from gas to liquid.

There are people who propose using high-density propellants for SSTOs.
But they are willing to go farther than you do:  the numbers say that
high-density propellants are better even if they are less energetic.
Larger mass ratios are necessary -- perhaps 20:1 rather than 10:1 -- but
they are also easier to achieve.  The first stage of the Titan II, built
in the early 1960s, has the mass ratio and engine performance to be an
(expendable) SSTO, thanks to dense propellants.

>...Another advantage
>of dense propellants is that you can build highly pressurized fuel and
>oxidizer tanks without losing too much weight. Combined with an engine
>that is effective even at low pressure ratios, like an aerospike engine,
>you might not need any turbopumps anymore...

Unfortunately, *this* doesn't work.  Mitch Burnside Clapp tried to put
such a design together about three years ago, using about the densest good
propellant combination available (kerosene and hydrogen peroxide).  The
back-of-the-envelope design sketch looked feasible, but when Mitch took a
hard look at issues like tank mass and pressurization (prodded by Bruce
Dunn, who's a big pressure-feed enthusiast), the conclusion was that it
wouldn't work.  The tanks still weigh too much.

>...Does anybody know wether SiH4 has been tested as a rocket fuel?

Yes, although not recently, and I believe the results weren't wonderful.

>Any information about exotic
>chemical propellants, like LiH/O2,SiH4/O2 or Be/O2 would be welcome.

The highest measured Isp ever achieved with chemical propellants was
with LH2/LF2/LLi (yes, liquid lithium), at 542s in vacuum.  Unfortunately,
it has all the problems of liquid fluorine, all the problems of liquid
lithium (which is not only hot but fiercely corrosive), and all the
problems of liquid hydrogen.  Pity.

Substituting oxygen for fluorine doesn't work nearly as well; the lithium
doesn't buy you a lot without the fluorine.

LH2/LOX/Be looks like a winner on paper, but not in practice.  It is
terribly expensive, horribly poisonous, and worst of all, it doesn't work
very well.  Almost any combination containing beryllium seems guaranteed
to have terrible combustion efficiency -- the beryllium just doesn't want
to burn -- and BeO has a sky-high melting point, just what isn't wanted
for efficient energy conversion.  Even a large dose of LH2 can't save
this one.
Space will not be opened by always                 |       Henry Spencer
leaving it to another generation.   --Bill Gaubatz |

From: Henry Spencer <>
Subject: Re: Advantages of Hydrogen vs. Hydrocarbons (for SSTOs)
Date: Thu, 30 Jan 1997 17:15:02 GMT

In article <5cl5tt$>,
Marcus Lindroos INF <> wrote:
>One of the assumptions about hydrocarbon LVs has always confused me. 
>Namely, that any tank capable of holding _x_ kilograms of hydrogen
>apparently can be expected to carry a far heavier fuel load of dense
>hydrocarbons without any weight-adding strengthening of the structure.

Yes, that is basically correct.  (In fact, the tank mass may well go down 
a bit, because the hydrocarbons don't need the insulation that hydrogen 

>In other words, pressure loads drive the tank loads? But you also
>have hydraulic loads from the weight of the liquid propellant 
>plus acceleration at up to 3G or so. From my limited understanding
>of the problem, this would be expected to be far higher than the 
>pressure load...

No -- do the numbers.  If we assume a 10m column of propellant with the
density of kerosene, at 3G the added pressure at the tank base from the
hydraulic load is about 35psi.  You will probably have to pressurize
your LH2 tank to at least that range just to suppress cavitation in the
pumps, which is much less of a problem with kerosene.

Only when you start talking about very large vehicles do inertia loads
dominate the tank structure.

>you will need more engine thrust as well -- producing yet more loads on 
>the structure as well as noise during liftoff. Engine mass savings for
>the kerosene vehicle do not appear to be that big, unless those Russian
>high-performance engines can produce a *far* higher T/W than 100-110 or

The Russian engines are up around 125, as I recall.  100-110 is what US
engines reached 35 years ago.

>For ELVs, the extremely high Isp of hydrogen is regarded as more important
>than the bulk density of kerosene during most of the ascent profile, 
>except for liftoff and initial ascent when thrust is more important than
>a high exhaust velocity. Do we have serious reasons to believe RLVs will
>be different, in this respect?

Yes -- note that almost all those hydrogen ELVs are *upper stages*.  For
an upper stage, there is one real advantage for hydrogen:  lower gross
mass means smaller lower stages. 

Note also that the classical comparisons of Isp vs. bulk density, which
most current orthodox thinking is still based on, assumed that the major
advantage of bulk density was lower air drag.  They assumed that engine
T/W would be equal, that tank mass scaled with propellant mass, and that
hydrogen would incur no significant penalties in auxiliary systems -- all
of which are now known to be false.

The point of the argument is not that RLVs are different from expendables,
but that orthodox thinking greatly exaggerates the advantages of hydrogen
even for expendables, because it is ultimately based on naive analyses
done many years ago using incorrect assumptions.
"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: Advantages of The Expander Cycle Engine
Date: Thu, 30 Jan 1997 16:34:00 GMT

In article <5cbbba$>,
Richard Varvill  <> wrote:
>I made no such assumption.  What I meant (although obviously didn't 
>state clearly enough) was taking into account the differences in 
>achievable mass ratio between the two propellant combinations and their 
>specific impulses, the hydrogen vehicle is able to achieve a higher 
>deltaV than the kerosene one.  This is fact, not a matter for 

Actually, it's a matter of considerable debate, because the assumptions
behind it are questionable.  As a case in point, the theoretical delta-V
of the Titan II first stage (using dense fuels, although as it happens,
not kerosene) is right up in the range achieved by the best hydrogen
stages.  The empirical evidence for your "fact" seems weak.

>The differences in required deltaV will be very small for,both vehicles 
>assuming that they are both starting from the ground.  Black Horse 
>achieves a useful reduction in the deltaV by virtue of the air-air 

No, the numbers Mitch came up with were for an apples-to-apples comparison,
*not* assuming that one system takes off from the ground and the other is
air-refuelled.  Gravity losses are significantly smaller at lower Isps,
because more of the vehicle mass is burned off earlier, and accelerations
are thus higher (or the vehicle's acceleration limit is reached earlier),
and the more rapid climb means a quicker transition to horizontal flight
free of gravity losses.

Mitch's numbers were 29050ft/s to orbit, using peroxide/kerosene, for
conditions in which a NASA LOX/LH2 design needed 31000.  Same orbit, same
launch, same initial acceleration, same G limit.  The calculation included
the lower drag of the denser vehicle, which helped a little.

This makes enough of a difference in the mass ratio needed for dense fuels
to bring the gross liftoff mass down into the same range as a LOX/LH2
"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: Why Slush Hydrogen?
Date: Fri, 3 May 1996 13:32:50 GMT

In article <4m6o4n$> (Pat) writes:
>even a 5% increase in fuel density translates to a larger payload.
>That 5% is almost pure profit.

Unfortunately, in the case of slush hydrogen, it comes with enough added
complications that the gain in density probably isn't worth it.  Handling
a mixture of liquid and solid isn't simple.  Even something as mundane as
a fuel gauge is difficult to build, because you care about the percentage
of solids as well as the total volume. 
Americans proved to be more bureaucratic           |       Henry Spencer
than I ever thought.  --Valery Ryumin, RKK Energia |

From: (Henry Spencer)
Subject: Re: Optimal LOX/LH2 mixture ratio
Date: Thu, 28 Oct 1999 18:32:42 GMT

In article <>,
Jukka Neppius  <> wrote:
>} >Variation of same idea is to use kerosene & LOX at start and switch to
>} >LH2 later.  Did somebody study this?
>} ...In general, the extra complexity does
>} not look to be worth the trouble.  Depending on what figures of merit you
>} use for evaluation, the optimum percentage of LH2 generally turns out to
>} be either 0% or 100%.  (If you use *sensible* figures of merit -- none
>} of this gross-liftoff-mass nonsense -- it's generally 0%.)
>SSTO needs many engines because large range of throttling needed.  I
>don't understand how it makes system more complex if some of those
>engines use different fuel.

Extra (insulated) tank with extra structural complications, more plumbing,
two different kinds of engine to develop and maintain, boiloff problems in
pre-launch phase, etc.  The complexity isn't huge, but then the benefits
aren't either.  Hydrogen's apparent performance advantage is much reduced
by more careful evaluation which includes its various problems.

>Large gross-liftoff-mass means powerful engines which are lifted to
>orbit (in SSTO), just like empty fuel tanks.  Is an empty fuel tank
>really much heavier than an engine needed to lift it when full?

It can be, yes:  the tanks are very large, they have to be insulated, the
plumbing has to be insulated too, complex chilldown arrangements are
needed, and to cap it off, the thrust:weight ratio of LOX/kerosene engines
is considerably better than that of LOX/LH2 engines.  So it's heavy tanks
plus heavy plumbing vs. a need for more thrust with a technology which
delivers high thrust more easily.

Compare the first stage of the Titan II with the second stage of the
Saturn V.  Both have, on paper, the mass ratio and Isp to deliver payloads
to orbit as expendable SSTOs.  Except that the S-II doesn't really have
quite enough thrust, and its engines are optimized for high altitude
operation and can't run properly at sea level.  The Titan II first stage,
on the other hand, runs just fine at sea level, and it has *too much*
thrust for an efficient trajectory -- its engines would have to be cut
down somewhat to make it work as an SSTO.  The S-II's development was a
memorable ordeal with heroic efforts at weight reduction.  The Titan II
development was straightforward, even though, as the smaller stage with
supposedly-inferior propellants, it theoretically should have had much
more difficulty achieving the same performance level.
The space program reminds me        |  Henry Spencer
of a government agency.  -Jim Baen  |      (aka

From: (Henry Spencer)
Subject: Re: Optimal LOX/LH2 mixture ratio
Date: Sun, 31 Oct 1999 02:55:03 GMT

In article <>, Paul DeMone  <> wrote:
>> the thrust:weight ratio of LOX/kerosene engines
>> is considerably better than that of LOX/LH2 engines.
>  Is this due to the ultra low density of LH2, the
>  higher temperature gradiant across engine components,
>  or something else?

It's a bunch of things.  One big factor is that pumps pump volume, not
mass, so the pumps for a bulky fluid are themselves bulky.  The same is
true, to a lesser extent, of the plumbing in general.  And bear in mind
that much of this hardware is pressure vessels to some degree, so more
volume equals more mass.

A lesser, but not unimportant, issue is that hydrogen engines are more
complex.  Most hydrogen plumbing needs to be insulated, often with vacuum
jackets, and often the engine needs provisions to prechill the hydrogen
plumbing so that the sudden start of LH2 flow won't cause flash boiling
(LH2 boils very easily).  All of this means more hardware, all of which
weighs something, and sometimes fewer options in design, which means you
have to pick something heavier.

A further complication is that hydrogen causes embrittlement in some
alloys.  Not a huge problem, but again, it limits your choices and can
force you to do suboptimal things.  The extremely low temperature likewise
imposes limitations.

>  Can you give some representative figures for thrust
>  to weight ratios for the two classes of engines?

75:1 is very good for a hydrogen engine, even today.  The LOX/kerosene
engines of the Saturn IB (first flight 1966) did over 100:1, and some of
the modern Russian engines reportedly approach 150:1 (although one has to
be careful about definitions, some of those Russian numbers apparently
don't include everything that Western designers would include in the
engine mass).

Gerald Nordley once told me that when you plot rocket-engine *power* per
kilogram, it's almost constant except for some SDIO experimental designs.
And jet power is proportional to thrust times Isp, so it's not surprising
that LOX/LH2, whose Isp is about 1.5x LOX/kerosene's, has about a 1.5x
engine mass penalty.
The space program reminds me        |  Henry Spencer
of a government agency.  -Jim Baen  |      (aka

From: (GCHudson)
Subject: Re: X-33 first flight?
Date: 9 Dec 1998 05:16:44 GMT

Jim Davis wrote:


>If one accepts the premise of the dense fuel advocates that vehicle
>empty mass is determined by propellant volume and not propellant weight
>both Vehicle 1 and Vehicle 2 must have identical empty weights. But
>since Vehicle 2 (the kerosine vehicle) has a higher orbiting mass its
>*payload* mass must be larger.


Jim, that is not a premise but about as certain as a fact may be.  Tank mass is
sized by propellant volumes so long as engines require any NPSH, propellants
have any vapor pressure and/or any ullage pressure is used to offset aero and
other loads.  In the real world, none of these may be neglected and thus all
tanks have some slight pressure.  If this is the case, the tank mass will be
determined by the volume of the propellants.  After analyzing dozens of
vehicles (both existing and viewgraph) I have never found ground launched
lox-hydrogen vehicles to be superior to lox-kerosene.  Air-launched is another
matter, for obvious reasons.

Gary C. Hudson, CEO
Rotary Rocket Company

From: (GCHudson)
Subject: Re: X-33 first flight?
Date: 9 Dec 1998 15:38:41 GMT

Allen Thomson wrote:

>>Jim, that is not a premise but about as certain as a fact may be.  Tank mass
>>sized by propellant volumes so long as engines require any NPSH, propellants
>                                                           ^^^^
>                                                          Whazzat?

Pump fed engines require some Net Positive Suction Head for the turbopumps to
function.  This means the tanks must have some pressurization in order to avoid
cavitation in the pump inlet.

Gary C. Hudson, CEO
Rotary Rocket Company

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