From: email@example.com (Henry Spencer)
Subject: SSTO delta-V and dense fuels
Date: Thu, 2 May 1996 23:49:51 GMT
First little bit of interesting news from Space Access 96, now that
I'm back and have assorted crises here more or less under control...
Those with good memories will recall that some SSTO supporters have been
advocating use of dense propellant combinations like H2O2/kerosene or
LOX/propane, instead of the orthodox LOX/hydrogen. The Isp is lower, and
hence you need a higher mass ratio, but the much greater density makes the
mass ratio easier to achieve. When examined with sophisticated scaling
models, rather than mindless "fixed fraction of propellant mass" ones, the
dry mass goes *down* -- despite an increase in gross mass -- because the
vehicle is smaller.
Well, Mitch Burnside Clapp has done it again. :-) He's found a big flaw
in a major assumption of the standard argument, and now dense fuels look
The incorrect assumption is that the total delta-V to reach orbit is the
same for all fuels. It's not. Dense fuels need less. Substantially less.
Consider two SSTOs, one LOX/LH2 and one H2O2/kerosene (I like LOX/propane
myself, but H2O2/kerosene is Mitch's favorite, and it's his discovery...),
with the same GLOM (gross liftoff mass), the same engine thrust (and so
same initial acceleration), and no requirement for G-limiting. Draw a
graph of mass vs. time for both.
Assume for the moment that they have the same total burn time. The curves
(well, lines) start from the same point. The H2O2/kerosene one has to get
rid of more mass, so to reach its final mass in the same amount of time,
the slope of its mass line must be steeper.
Wait a minute. A steeper mass line means that at any time after liftoff,
the H2O2/kerosene SSTO has lower mass than the LOX/LH2 one, and since they
have the same thrust... the H2O2/kerosene SSTO is accelerating faster. If
they have the same total delta-V requirement, that last assumption must be
wrong: the H2O2/kerosene burn time is shorter.
But... the biggest penalty on top of the theoretical delta-V is gravity
losses, and gravity losses are a function of burn time! The H2O2/kerosene
SSTO is accelerating faster, so it has lower gravity losses, and needs
less total delta-V. Moreover, that makes its burn time still shorter, and
its mass line still steeper, so the difference in acceleration is even
larger than it first seems.
Adding G-limiting, which is a practical necessity, changes the details
but not the overall result: the dense-propellant SSTO loses mass faster,
accelerates faster before G-limiting, and so has lower gravity losses.
The bottom line, when all this converges -- including a small gain from
lower drag on a more compact vehicle, and a very small bonus from lower
drag making the acceleration still higher -- is that a standard orthodox
NASA LOX/LH2 SSTO needs 31000ft/s to reach the space-station orbit, and an
H2O2/kerosene SSTO needs only 29050ft/s.
(In fact, the explanation came after the numbers -- when good trajectory
simulations kept coming out with lower delta-Vs for H2O2/kerosene, Mitch
decided he had to understand what was going on.)
Now, consider. The H2O2/kerosene SSTO is operating in a very steep part
of the mass-ratio curve. A 6% saving in delta-V is *not* trivial. For
engines with a vacuum Isp of 320, the required mass ratio drops from 20 to
16. Given the aforementioned sophisticated scaling models, at this mass
ratio, the H2O2/kerosene SSTO's payload at the same GLOM is now equal to
that of the LOX/LH2 design.
So the dense-fuel SSTO has lower dry mass, smaller vehicle size, cheaper
and easier-to-handle propellants, and now suffers no GLOM penalty... Just
what was the advantage of LOX/LH2 supposed to be again?
Americans proved to be more bureaucratic | Henry Spencer
than I ever thought. --Valery Ryumin, RKK Energia | firstname.lastname@example.org
From: Henry Spencer <email@example.com>
Subject: Re: SSTO delta-V and dense fuels
Date: Mon, 13 May 1996 18:49:37 GMT
In article <firstname.lastname@example.org> Jordin Kare <email@example.com> writes:
>>The incorrect assumption is that the total delta-V to reach orbit is the
>>same for all fuels. It's not. Dense fuels need less...
>With all due credit to Mitch, the reduced gravity loss of lower Isp
>propellants is not news. What is surprising is that the advantage is
>as large as ~2000 feet per second.
It's been clear all along that there were minor advantages, both from
unloading more mass at low altitude and from reduced drag of a smaller
vehicle. Mitch didn't just do the numbers more carefully; he uncovered a
whole new effect -- the higher acceleration -- which (as far as I know)
hadn't been noticed before, and which is much more important than the
previously-known effects. Hence the big number.
>This may depend strongly on some of the model parameters... Low-Isp
>systems probably optimize at a higher initial thrust-to-weight...
Actually, at least one past study has concluded the opposite, that
they optimize at lower T/W. (See Martin&Manski in the Nov/Dec 1991
Journal of Propulsion & Power.) I wonder if this wasn't the result
of, in effect, optimizing for constant gravity losses.
>It would be interesting to "tweak" the models used...
Yes indeed. It's a very interesting result that deserves exploration.
Unix was a breakthrough. | Henry Spencer
Windows 95 is more like a smash-and-grab. | firstname.lastname@example.org
From: Henry Spencer <email@example.com>
Subject: Re: SSTO delta-V and dense fuels
Date: Tue, 14 May 1996 03:24:16 GMT
In article <firstname.lastname@example.org> email@example.com (Marcus Lindroos INF) writes:
>...It seems to me as if dense,
>low-Isp fuels become more competitive as the required delta-V
>decreases, and if the vehicle is very small. Black Horse is
>an excellent example. I suspect that a kerosene/peroxide Delta Clipper
>(=large SSTO, difficult delta-V requirements) would be a lot harder.
Actually, among Mitch's pre-Black-Horse work was a scaling study of what
would have to change if you revised DC to burn peroxide/kerosene. The
gross liftoff mass went up; the dry mass went down; it looked feasible.
(And that was before the new delta-V results.)
>Now, if you keep the same GLOW mass but decrease the exhaust velocity
>to 3.2km/s (=high performance JP-5/H2O2 engine), the mass of the empty
>vehicle would have to decrease to 13t...
Remember that almost everything in the dry mass scales with propellant
*volume* or some function thereof. Denser propellants do indeed yield
considerably lower dry mass.
Unix was a breakthrough. | Henry Spencer
Windows 95 is more like a smash-and-grab. | firstname.lastname@example.org
From: email@example.com (burnside)
Subject: A LO2/kerosene SSTO rocket design, w/o AOL
Date: 2 Feb 1997 15:11:33 GMT
A LO2/Kerosene SSTO Rocket Design (long)
Mitchell Burnside Clapp
(view with a fixed pitch font such as courier or monaco)
The NASA Access to Space LO2/hydrogen single stage to orbit
rocket was examined, and the configuration reaccomplished
with LO2/kerosene as the propellants. Four major changes
were made in assumptions. First, the aerodynamic
configuration was changed from a wing with winglets to a
swept wing with vertical tail. The delta-V for ascent
was as a result recalculated, yielding a lower value due to
different values for drag and gravity losses. The engines
were changed to LO2/kerosene burning NK-33 engines, which
have a much lower Isp than SSME-type engines used in the
access to space study, but also have a much higher
thrust-to-weight ratio. The orbital maneuvering system on
the Access to Space Vehicle was replaced with a pump-fed
system based on the D-58 engine used for that purpose now
on Proton stage 4 and Buran. Finally, the wing of the
vehicle was allowed to be wet with fuel, which is a
reasonable practice with kerosene but more controversial
with oxygen or hydrogen. Additionally, in order to reduce
the technology development needed, the unit weights of the
tankage were allowed to increase by 17 percent.
After the design was closed and all the weights
recalculated, the empty weight of the LO2/kerosene vehicle
was 35.6% lighter than its hydrogen fuelled counterpart.
NASA completed a study in 1993 called Access to Space, the
purpose of which was to consider what sort of vehicle
should be operated to meet civil space needs in the future.
The study had three teams to evaluate three different broad
categories of options. The Option 3 team eventually settled
on a configuration called the SSTO/R. This vehicle was a
LO2/hydrogen vertical takeoff horizontal landing rocket.
The mission of the Access to Space vehicle was to place a
25,000 pound payload in a 220 n.mi. orbit inclined at 51.6
degrees. The vehicle had a gross liftoff weight of about
2.35 million pounds. The thrust at liftoff was 2.95 million
pounds, for a takeoff thrust to weight ratio of 1.2. The
empty weight of the vehicle was 222,582 pounds, and the
propellant mass fraction (defined here as
[GLOW-empty]/GLOW) was 90.5%.
Main power for this vehicle was provided by seven SSME
derivative engines, with the nozzle expansion ratio reduced
to 50. This resulted in an Isp reduction from 454 to 447.3
seconds. Each engine weighed 6,790 lbs, for an engine sea
level thrust to weight ratio of 62.
Aerodynamically the vehicle was fairly squat, with a
fineness ratio (length:diameter) of 5. The overall length
of the vehicle was 173 feet and its diameter was 34.6 feet.
It had a single main wing (dry of all propellants) of about
4,200 square feet total area, augmented by winglets for
directional control at reentry. The landing wing loading
was about 60 lb/ft2. The oxygen tank was in the nose
section. The payload was mounted transversely between the
oxygen and hydrogen tanks, and was 15 feet in diameter and
30 feet long.
This design exercise was among the most thorough ever
conducted of a single stage to orbit LO2/LH2 VTHL rocket.
It was probably the single greatest factor in convincing
the space agency that single stage to orbit flight was
feasible and practical, to borrow from the title of Ivan
Bekey's paper of the same name.
A LO2/kerosene alternative
A number of people have been asserting for some time that
higher propellant mass fractions available from dense
propellants may make single stage to orbit possible with
those propellants also. The historical examples of the
extraordinary mass fractions of the Titan II first stage,
the Atlas, and the Saturn first stage are all persuasive.
Further, denser propellants lead to higher engine thrust to
weight ratios, for perfectly understandable hydraulic
It has not usually been observed that higher density also
leads to significant reductions in required delta-v.
There are two major reasons that this is so. First, the
reduction in volume leads to a smaller frontal area and
lower drag losses. The second, and more significant, reason
is that the gravity losses are also reduced. This is because
the mass of the vehicle declines more rapidly from its
initial value. The gravity losses are proportional to the
mass of the vehicle at any given time, and hence the
vehicle reaches its limit acceleration speed faster.
NASA itself has implicitly recognized this effect. When the
Access to Space Option 3 team examined tripropellant
vehicles, the delta-v to orbit derived from their work was
29,127 ft/sec, for precisely the reasons described in the
previous paragraph. This compares to a delta-v of 30,146
ft/s for the hydrogen-only baseline, as reported in a
briefing by David Anderson of NASA MSFC dated 6 October
1993. To be clear, these delta-v numbers include the back
pressure losses, so that no "trajectory averaged Isp"
number is used. They did not, however, report any results
for kerosene-only configurations.
To come to a more thorough understanding of the issues
involved in SSTO design, I have used the same methodology
as the Access to Space team to develop compatible numbers
for a LO2/kerosene SSTO. There are four major changes in
basic assumption between the two approaches, which I will
identify and justify here:
1: The ascent delta-v for the LO2/kerosene vehicle is
29,100 ft/sec, rather than 29,970 ft/sec. The reason for
this is argued above, but I ran POST to verify this value,
just to be sure. The target orbit is the same: 220 n.mi.
circular at 51.6 degrees inclination. The detailed weights
I have for the NASA vehicle are based on a delta-v of
29,970 ft/sec rather than the 30,146 ft/sec reported in
Anderson's work, but I prefer to use the values more
favourable to the hydrogen case to be conservative. The
optimum value of thrust to weight ratio turns out to be
slightly less than the hydrogen vehicle: 1.15 instead of
2: The aerodynamic configuration is that of Boeing's RASV.
Without arguing whether this is optimal, the fineness ratio
of 8.27 and large wing lead to a much more airplane-like
layout, better glide and crossrange performance, and
reduced risk. The single vertical tail is simpler and safer
than winglets as well. Extensive analysis has justified the
reentry characterisitics of this aircraft. The wing is
assumed to be wet with the kerosene fuel, as is common on
most aircraft. The fuel is also present in the wing
carry-through box. The payload is carried over the wing
box, and the oxidizer tank is over the wing. This avoids
the need for an intertank, which in the NASA Access to
Space design is nearly 6,600 pounds.
3. The main propulsion system is the NK-33. The engine has
a sea level thrust of 339,416 lbs, a weight of 2,725 lbs
with gimbal, and a vacuum Isp of 331 seconds. Furthermore,
it requires a kerosene inlet pressure of only 2 psi
absolute, which dramatically reduces the pressure required
in the wing tank. It also operates with a LO2 pressure at
the inlet of only 32 psi. The comparable values for the
SSME are about 50 psi for both propellants. This will have
a substantial effect on the pressurization system weight.
4. The OMS weight is based on the D-58 engine. This engine
is used for the Buran OMS system and the Proton stage 4. As
heavy as it is the Isp is an impressive 354 seconds. NASA's
vehicle used a pressure fed OMS, which is a sensible design
choice if you're stuck with hydrogen and you wish to
minimize the number of fluids aboard the vehicle. But
because both oxygen and kerosene are space-storable, there
is no reason to burden the design with a heavy pressure fed
Using the same methodology for calculating masses, and
accepting the subsystems masses as given in the Access to
Space vehicle, a redesign with oxygen and kerosene was
accomplished. The results appear in Table 1.
Table 1: Access to Space vehicle and LO2/kerosene
Name O2/H2 LO2/RP
Wing 11,465 11,893 lb
Tail 1,577 1,636 lb
Body 64,748 33,741 lb
Fuel tank 30,668 - lb
Oxygen tank 13,273 17,271 lb
Basic Structure 14,610 10,274 lb
Secondary Structure 6,197 6,197 lb
Thermal Protection 31,098 21,238 lb
Undercarriage, aux. sys 7,548 5,097 lb
Propulsion, Main 63,634 36,426 lb
Propulsion, RCS 3,627 1,234 lb
Propulsion, OMS 2,280 823 lb
Prime Power 2,339 2,339 lb
Power conversion & dist. 5,830 5,830 lb
Control Surface Actuation 1,549 1,549 lb
Avionics 1,314 1,314 lb
Environmental Control 2,457 2,457 lb
Margin 23,116 16,105 lb
Empty Weight 222,582 141,682 lb
Payload 25,000 25,000 lb
Residual Fluids 2,264 1,911 lb
OMS and RCS 1,614 1,261 lb
Subsystems 650 650 lb
Reserves 7,215 8,895 lb
Ascent 5,699 7,587 lb
OMS 679 541 lb
RCS 837 767 lb
Inflight losses 13,254 17,445 lb
Ascent Residuals 10,984 15,175 lb
Fuel Cell Reactants 1,612 1,612 lb
Evaporator water supply 658 658 lb
Propellant, main 2,054,612 3,034,972 lb
Fuel 293,604 843,048 lb
Oxygen 1,761,008 2,191,924 lb
Propellant, RCS 2,814 2,556 lb
Orbital 2,051 1,756 lb
Entry 763 800 lb
Propellant, OMS 19,357 15,452 lb
GLOW 2,347,098 3,246,156 lb
Inserted Weight 292,486 211,185 lb
Pre-OMS weight 271,482 186,152 lb
Pre-entry Weight 252,125 170,700 lb
Landed Weight 251,362 169,900 lb
Empty weight 222,582 141,682 lb
Sea Level Thrust 2,816,518 3,733,080 lb
Percent margin 11.6% 12.8%
Assumed Isp(vac) 447.3 331.0 s
Ascent Delta-V 29,970 29,100 ft/s
OMS delta-V 1,065 987 ft/s
RCS delta-V 108 107 ft/s
Deorbit Delta-V 44 53 ft/s
Reserves 0.28% 0.25% lb/lb
Residuals 0.53% 0.50% lb/lb
Wing Parameter 4.56% 7.00% lb/lb
TPS parameter 12.37% 12.50% lb/lb
Undercarriage parameter 3.00% 3.00% lb/lb
Wing Reference Area 4,189 5,528 ft2
Density of fuel 4.4 50.5 lb/ft3
Density of oxygen 71.2 71.2 lb/ft3
Volume of fuel 66,276 16,694 ft3
Volume of oxygen 24,733 30,785 ft3
Fuel tank parameter 0.42 - lb/ft3
Oxygen tank parameter 0.48 0.56 lb/ft3
Some discussion of the results and justification is in
The wing is about 40 percent heavier as a percentage of
landed weight than for the hydrogen fueled baseline. When
considered as a tank, it is about 60 percent heavier for
the volume of fuel it encloses. Its weight per exposed area
is about the same and the wing loading is half at landing.
No benefit is taken explicitly for the lack of a
requirement for kerosene tank cryogenic insulation.
The tail is assumed to have the same proportion of wing
weight for both cases. This is conservative for the
kerosene wehicle because its single vertical tail is
structurally more efficient.
The body of the kerosene vehicle has three components. The
oxidizer tank has an increased unit weight of about 17
percent. This is done in order to avoid the need for
aluminum-lithium, which was assumed in the Access to Space
vehicle. The basic structure group is unchanged, except
that the intertank is deleted and the thrust structure is
increased in proportion to the change in thrust level.
The secondary structure group is mostly payload support
related, and was not changed.
The thermal protection group is in both cases about 12.5%
of the entry weight. This works out to 1.107 lbs/ft2 of
wetted area for the kerosene vehicle, which is common to
many SSTO designs.
The undercarriage group is 3% of landed weight for both
vehicles. There is no benefit taken for reductions in gear
loads for the kerosene vehicle due to lower landing speed
and lower glide angle at landing.
The main propulsion group includes engines, base mounted
heat shield, and pressurization/feed weights. The engines
are far lighter for their thrust than SSME derivatives. The
pressurization weights are reduced in proportion to the
pressurized volume for the kerosene vehicle. No benefit is
taken for reduced tank pressure.
Here is as good a place as any to point out the erroneous
assertion that increased hydrostatic pressure is going to
lead to increased tankage weights. There is no requirement
for a particular ullage pressure except for the need to
keep the propellants liquid. It is the pressure at the base
of the fluid column rather than the top of the column that
is of engineering interest. The column of fluid exerts a
hydrostatic load on the base of the tank, but this load
does not typically exceed the much more adverse requirement
for engine inlet pressurization. For the kerosene vehicle,
the hydrostatic load at the base of the oxygen tank is 49
psi, which is compatible with the pressures normally seen
in oxygen tanks for rocket use. The load declines after
launch because the weight goes down faster than the
acceleration goes up.
The bottom line here is that dense propellants may require
you to alter a tank's pressurization schedule, but not to
overdesign the entire tank. Structures are sized by loads
and tankage for rockets is sized principally by volume, and if
the vehicle is small, by minimum gauge considerations.
This is not completely true for wet wings, however, as
discussed previously. In this particular example, there is
no need for high pressure in the wing tank either, because
of the low inlet pressure required by the NK-33.
The OMS group is the only other major change, as discussed
above. The reliable D-58 engine has been performing space
starts for decades and will serve well here. The
acceleration available from the OMS is about 0.12 g, which
All the other weights are pushed straight across for the
most part. A brief inspection suggests that this is very
conservative. Control surface actuation requirements are
certainly less, electrical power requirements less, much
better fuel cells available than the phosporic acid type
assumed here, and reduced need for environmental control.
Nonetheless, rather than dispute any of these values it is
easier simply to accept them.
The margin is applied to all weight items at 15% execpt for
the engine group at 7.5%. The justification for this is that
the main and OMS engine weights are known to high accuracy.
The vehicle has an overall length of 1955 inches, and a
diameter of 236.4 inches. The wing has a leading edge sweep
of 55.5 degrees and a trailing edge sweep of -4.5 degrees.
Its reference area is 5,632 square feet, of which 3,992
square feet is exposed. The wing encloses 16,694 ft3 of
fuel, with a further 5% ullage. The carry-through is also
wet with fuel. The wing span is 1293 inches, and the taper
ratio is 0.13.
The payload bay has a maximum width and height of 15 feet.
It sits on top of the wing carry through box. The thrust
structure from the engines passes through and around the
payload bay to the forward LO2 tank. The payload bay is 30
feet in length. It has a pair of doors, the aft edge of
which is just forward of the vertical tail leading edge.
The engine section encloses 11 NK-33 engines, with a 4 - 3
- 4 layout. The engines are each 12.5 feet long, and
additional structure and subsystems take up another 6.5
The oxygen tank comprises the forward fuselage, which
encloses 30,785 ft3 of oxygen, with a further 5% ullage.
The length of the tank is about 100 feet. The ventral
surface of the tank is moderately flattened as it moves
aft, to fair smoothly with the wing lower surface. This
flattening reduces its length by about 5% with respect to a
strictly cylindrical layout. The aft edge of the oxygen tank
is about even with the forward payload bay bulkhead. A
compartment of about 13.9 feet provides room for some
subsystems and a potential cockpit in future versions.
The methods of the NASA Access to Space study were used to
design a single stage to orbit vehicle using existing
LO2/kerosene engines. An inspection of the final results
shows that the vehicle weighs about 36.5% less than its
hydrogen counterpart, with reductions in required
technology level and off the shelf engines. The center of
mass of the vehicle is about 61% of body length rather than
68% for the Access to Space vehicle, which should improve
control during reentry. The landing safety is considerably
improved by lower landing speed and better glide ratio.
Structural margins are greater overall. The vehicle
designed here appears to be superior in every respect:
smaller, lighter, lower required technology, improved
safety, and almost certainly lower development and