A Kerosene-Fueled X-33 as a Single Stage to Orbit Vehicle.

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Table of Contents.
II.)Lightweight propellant tanks.
III.)Kerosene fuel and engines for the X-33/Venture star.
IVa.)Aerodynamic lift applied to ascent to orbit.
b.)Estimation of fuel saving using lift.
V.)Kerosene fueled VentureStar payload to orbit.

I.) A debate among those questing for the Holy Grail of a reusable,
single-stage-to-orbit vehicle is whether it should be powered by hydrogen or a
dense hydrocarbon such as kerosene. Most concepts for such a vehicle centered
on hydrogen, since a hydrogen/LOX combination provides a higher Isp. However,
some have argued that dense fuels should be used since they take up less
volume (equivalently more fuel mass can be carried in the same sized tank) so
they incur less air drag and also since the largest hydrocarbon engines
produce greater thrust they can get to the desired altitude more quickly so
they also incur lower gravity drag loss.
Another key fact is that for dense fuels the ratio of propellant mass to tank
mass is higher, i.e., you need less tank mass for the same mass of propellant.
This fact is explored in this report:

Single Stage To Orbit Mass Budgets Derived From Propellant Density and
Specific Impulse.
John C. Whitehead
32nd AIAA/ASME/SAE/ASEE Joint Propulsion ConferenceLake Buena Vista, FLJuly
1-3, 1996
http://www.osti.gov/bridge/servlets/pur ... 379977.pdf

Whitehead notes that the propellant mass to tank mass ratio for kerosene/LOX
is typically around 100 to 1, while for liquid hydrogen/LOX it's about 35 to
1, which would result in a significantly greater dry mass for the
hydrogen-fueled case just in tank weight alone. Based on calculations such as
these Whitehead concludes the best option for a SSTO would be to use
The case for the X-33/VentureStar is even worse because the unusual shape of
the tanks requires them to use more tank mass than a comparably sized
cylindrical tank. This is discussed here:

Space Access Update #91 2/7/00.
The Last Five Years: NASA Gets Handed The Ball, And Drops It.
"...part of L-M X-33's weight growth was the "multi-
lobed" propellant tanks growing considerably heavier than promised.
Neither Rockwell nor McDonnell-Douglas bid these; both used proven
circular-section tanks. X-33's graphite-epoxy "multi-lobed" liquid
hydrogen tanks have ended up over twice as heavy relative to the
weight of propellant carried as the Shuttle's 70's vintage aluminum
circular-section tanks - yet an X-33 tank still split open in test
last fall. Going over to aluminum will make the problem worse; X-
33's aluminum multi-lobed liquid oxygen tank is nearly four times as
heavy relative to the weight of propellant carried as Shuttle's
aluminum circular-section equivalent."

The X-33's twin liquid hydrogen tanks had a weight of 4,600 pounds each, and
the liquid oxygen tank a weight of 6,000 pounds, for total of 15,200 pounds
for the tanks:

Marshall Space Flight Center
Lockheed Martin Skunk Works
Sept. 28, 1999
X-33 Program in the Midst of Final Testing and Validation of Key Components.

The weight of the propellant carried by the X-33 was supposed to be 210,000
lb. So the propellant to tank mass ratio for the X-33 was only about 14 to
1(!). This would be a severe problem for the full-scale VentureStar. Its gross
lift off weight was supposed to be 2,186,000 lbs with a fuel weight of
1,929,000 lbs:

X-33 Advanced Technology Demonstrator.
http://teacherlink.ed.usu.edu/tlnasa/Ot ... trator.PDF

So the VentureStar would have a dry mass of 257,000 lbs. Since the same design
would be used for the VentureStar tanks as those of the X-33, the propellant
mass to tank mass ratio would also be 14 to 1, so the tank mass would be
138,000 lbs. But this means the empty tank mass alone would be over half of
the vehicle's dry weight (!)
It would have been extremely difficult for the VentureStar to have made orbit
with such a large weight penalty from the start. From all accounts the weight
problem with the tanks drove other problems such as the need to add larger
wings, increasing the weight problem further. NASA wound up canceling the
program when Lockheed couldn't deliver the working liquid hydrogen tanks even
at this excessive weight. However, rather than canceling the program I believe
the better course would have been to open up competition for coming up with
alternative, creative solutions for reducing the weight of the tanks. This
would also have resolved some of the problems with the vehicles weight growth.

II.) I have proposed one possibility for lightweighting the X-33 tanks on this

http://www.bautforum.com/space-explorat ... ost1495726

The idea would be to achieve the same lightweight tanks as cylindrical ones by
using multiple, small diameter, aluminum cylindrical tanks. You could get the
same volume by using varying lengths and diameters of the multiple cylinders
to fill up the volume taken up by the tanks. The cylinders would not have to
be especially small. In fact they could be at centimeter to millimeter
diameters, so would be of commonly used sizes for aluminum tubes and pipes.
The weight of the tanks could be brought down to the usual 35 to 1 ratio for
propellant to tank mass. Then the mass of the tanks on the X-33 would be
210,000 lbs/35 = 6,000 lbs, saving 9,200 lbs off the vehicle dry weight. This
would allow the hydrogen-fueled X-33 to achieve its original Mach 15 maximum
The same idea applied to the full-scale hydrogen-fueled VentureStar would
allow it to significantly increase its payload carrying capacity. At a 35 to 1
ratio of propellant mass to tank mass, the 1,929,000 lbs propellant mass would
require a mass of 1,929,000/35 = 55,000 lbs for the tanks, a saving of 83,000
lbs off the original tank mass. This could go to extra payload, so from 45,000
lbs max payload to 128,000 lbs max payload.
An analogous possibility might be to use a honeycombed structure for the
entire internal makeup of the tank. The X-33's carbon composite tank was to
have a honeycombed structure for the skin alone. Using a honeycomb structure
throughout the interior might result in a lighter tank in the same way as does
multiple cylinders throughout the interior.
Still another method might be to model the tanks standing vertically as
conical but with a flat front and back, and rounded sides. Then the problem
with the front and back naturally trying to balloon out to a circular cross
section might be solved by having supporting flat panels at regular intervals
within the interior. The X-33 composite tanks did have support arches to help
prevent the tanks from ballooning but these only went partially the way
through into the interior. You might get stronger a result by having these
panels go all the way through to the other side.
These would partition the tanks into portions. This could still work if you
had separate fuel lines, pressurizing gas lines, etc. for each of these
partitions and each got used in turn sequentially. A preliminary calculation
based on the deflection of flat plates under pressure shows with the tank made
of aluminum alloy and allowing deflection of the flat front and back to be
only of millimeters that the support panels might add only 10% to 20% to the
weight of the tanks, while getting similar propellant mass to tank mass ratio
as cylindrical tank. See this page for an online calculator of the deflection
of flat plates:

eFunda: Plate Calculator -- Simply supported rectangular plate with uniformly
distributed loading.
http://www.efunda.com/formulae/solid_me ... niform.cfm

Note you might not need to have a partitioned tank, with separate fuel lines,
etc., if the panels had openings to allow the fuel to pass through. These
would look analogous to the wing spars in aircraft wings that allow fuel to
pass through. You might have the panels be in a honeycomb form for high
strength at lightweight that still allowed the fuel to flow through the tank.
Or you might have separate beams with a spaces between them instead of solid
panels that allowed the fuel to pass through between the beams.
Another method is also related to the current design of having a honeycombed
skin for the composite hydrogen tanks. Supposed we filled these honeycombed
cells with fluid. It is known that pressurized tanks can provide great
compressive strength. This is in fact used to provide some of the structural
strength for the X-33 that would otherwise have to be provided by heavy
strengthening members. This idea would be to apply fluid filled honycombed
cells. However, what we need for our pressurized propellant tanks is *tensile
A possible way tensile strength could be provided would be to use the
Poisson's ratio of the honeycombed cells:

Poisson's ratio.

Poisson's ratio refers to the tendency of a material stretched in one
direction to shrink in length in an orthogonal direction. Most isotropic solid
materials have Poisson's ratio of about .3. However, the usual hexagonal
honeycombed structure, not being isotropic, can have Poisson's ratios in the
range of +1. This is mentioned in this article about non-standard honeycombed
structures that can even have negative Poisson ratios:

Chiral honeycomb.

However, note that from the formula for the volumetric change in the Wikipedia
Poisson's ratio page, a stretching of a material with a +1 Poisson's ratio
implies a *decrease* in volume; actually this is true for any case where the
Poisson's ratio is greater than +.5. Then fluid filled honeycombed cells would
resist the stretching of tensile strain by the resistance to volume
compression. This would be present with both gases and liquids. Gases are
lighter. However, they are highly compressible and it might take too large an
internal pressure in the cells to provide sufficient resistance, and so also
too thick cell walls to hold this pressure. Liquids are heavier but they are
highly non-compressible so could provide strong resistance to the volume
compression and thereby to the tensile strain.
Then for liquid hydrogen tanks we might use liquid hydrogen filled cells
within the skin of the tanks. Hydrogen is rather light compared to other
liquids at a density of only about 72 kg/m^3. This then could provide high
tensile strength at a much lower weight than typical solid wall tanks.
Kerosene and liquid oxygen would be used in the honeyombed cells for their
corresponding tanks, to keep the storage temperatures comparable. These are
heavier liquids than liquid hydrogen, approximately in the density range of
liquid water. Still these liquid filled honeycombed cells would provide much
lighter tanks than comparable solid wall tanks.

III.) Any of these methods might allow you to reduce the weight of the tanks
to be similar to that of cylindrical tanks and thus raise the payload to over
100,000 lbs. This would be for keeping the original hydrogen/LOX propellant.
However, in keeping with the analyses that show dense propellants would be
more appropriate for a SSTO vehicle I'll show that replacing the
hydrogen-fueled engines of the X-33/VentureStar with kerosene ones would allow
the X-33 to actually now become an *orbital* craft instead of just suborbital,
and the payload capacity of the VentureStar would increase to be comparable to
that proposed for Ares V.
The volume of the X-33 liquid hydrogen tanks was 29,000 gallons each and the
liquid oxygen tank, 20,000 gallons, for a total of 78,000 gallons volume for
propellant. This is 78,000gal*3.8 L/gal = 296,000 liters, 296 cubic meters.
How much mass of kerosene/LOX could we fit here if we used these as our
propellants? Typically the oxidizer to fuel ratio for kerosene/LOX engines is
in the range of 2.5 to 2.7 to 1. I'll take the O/F ratio as 2.7 to 1. The
density of kerosene is about 806 kg/m^3 and we can take the density of liquid
oxygen to be 1160 kg/m^3 when densified by subcooling:

Liquid Oxygen Propellant Densification Unit Ground Tested With a Large-Scale
Flight-Weight Tank for the X-33 Reusable Launch Vehicle.
http://www.grc.nasa.gov/WWW/RT/RT2001/5 ... omsik.html

These requirements of the propellants' total volume and densities, result in a
total propellant mass of 307,000 kg, with 83,000 kg in kerosene and 224,000 kg
in LOX. Kerosene/LOX tanks weigh typically 1/100th the propellant mass, so the
tank mass would be 3,070 kg. The current X-33 LH/LOX tanks weighed 15,200 lbs,
or 6,900 kg. So the empty weight of the X-33 is reduced from 63,000 lbs,
28,600 kg, to 28,600kg - 6,900kg + 3070kg = 24,800 kg.
How about the engines? The X-33 is to be reusable so you want to use reusable
kerosene engines. The RS-84 might be ideal when it is completed for the
full-scale VentureStar, but it turns out it's a bit too heavy for the X-33.
It would have a weight of about 15,000 lbs, 6,800 kg:


about the weight of the two aerospike engines currently on the X-33:

Bringing launch costs down to earth.
"Three federally funded projects are underway to develop new rocket engines
that can make it more affordable to send payloads into orbit."
http://www.memagazine.org/backissues/me ... aunch.html

With 307,000 kg kerosene/LOX fuel and 24,800 kg dry weight, the mass ratio
would be 13.4. According to the Astronautix page, the sea level Isp of the
RS-84 would be 301 s, and the vacuum 335 s. Take the average Isp as 320s. The
total Isp for a rocket to orbit including gravity and air drag losses is
usually taken to be about 9,200 m/s. Then an average exhaust velocity of 3200
m/s and mass ratio of 13.4 would give a total delta-v of 8,300 m/s. Even if
you add on the 462 m/s additional velocity you can get for free by launching
at the equator this would not be enough for orbit.
So for the X-33 I'll look at the cases of the lighter for its thrust NK-33,
used as a trio. Note that though not designed to be a reusable engine to make,
say, 100 flights, all liquid fuel rocket engines undergo extensive static
firings during testing so the NK-33 probably could make 5 to 10 flights before
needing to be replaced.The NK-33 is almost legendary for its thrust to weight
ratio of 136. According to the Astronautix page its weight is 1,222 kg , with
a sea level Isp of 297 sec and a vacuum Isp of 331:


I'll take the average Isp as 315 s. With three NK-33 engines the mass of the
X-33 becomes 21,700 kg, and the mass ratio becomes 15.15. Then with an average
Isp of 315 s, the total delta-v would be 8561 m/s and if you add on the 462
m/s additional equatorial velocity it's 9,023 m/s. Still slightly below the
delta-v typically given for orbit of 9,200 m/s.
However, it should be noted that the extra delta-v required beyond the 7,800
m/s orbital velocity is highly dependent on the vehicle and trajectory. Here's
a page that gives the gravity loss and air drag loss for some orbital rockets:

Drag: Loss in Ascent, Gain in Descent, and What It Means for Scalability.
Thursday 2008.01.10 by gravityloss
* Ariane A-44L: Gravity Loss: 1576 m/s Drag Loss: 135 m/s
* Atlas I: Gravity Loss: 1395 m/s Drag Loss: 110 m/s
* Delta 7925: Gravity Loss: 1150 m/s Drag Loss: 136 m/s
* Shuttle: Gravity Loss: 1222 m/s Drag Loss: 107 m/s
* Saturn V: Gravity Loss: 1534 m/s Drag Loss: 40 m/s (!!)
* Titan IV/Centaur: Gravity Loss: 1442 m/s Drag Loss: 156 m/s
http://gravityloss.wordpress.com/2008/0 ... alability/

Note that the gravity loss for the Delta 7925 is particularly small. As a
general principle the gravity loss can be minimized if you have a high thrust
vehicle that rapidly develops high vertical velocity sufficient to reach the
altitude for orbit. For then it can more quickly apply the horizontal thrust
required to achieve the 7,800 m/s orbital, tangential, velocity, there being
no gravity loss over the horizontal thrust portion. Note then the liftoff
thrust to liftoff weight ratio for the Delta 7925 is the relatively large 1.4;
in comparison for the Saturn V it was only 1.14. And note now also that the
X-33 with three NK-33 engines has total mass of 328,700 kg and total thrust of
4,530,600 N giving a liftoff thrust to liftoff weight ratio of 1.4. Then this
reconfigured X-33 would likely have comparable gravity drag loss as the Delta
7925. When you take into account the high thrust means it would rapidly reach
high altitude, implying the Isp would quickly get close to the vacuum Isp, the
average Isp over the trajectory is most likely closer to the 331 s vacuum Isp
than just 315 s, giving an actually higher achieved delta-v.

IV.a) It's capability of reaching orbit and possibly even with a small payload
could be increased with another additional factor. Among the questors for a
SSTO vehicle, the idea to use wings for a horizontal landing has been derided
because of the view they were just dead weight on ascent and you need to save
as much weight off the empty weight of the vehicle as possible to achieve
orbit. However, the key fact is that wings or a lifting body shape can reduce
the total delta-v for orbit by using aerodynamic lift to supply the force to
raise the vehicle to a large portion of the altitude of orbit rather than this
force being entirely supplied by the thrust of the engines. This fact is
discussed on page 4 of this report:

AIAA 2000-1045
A Multidisciplinary Performance Analysis of a Lifting-Body
Single-Stage-to-Orbit Vehicle.
Paul V. Tartabini, Roger A. Lepsch, J. J. Korte, Kathryn E. Wurster
38th Aerospace Sciences Meeting & Exhibit.
10-13 January 2000 / Reno, NV
"One feature of the VentureStar design that could
be exploited during ascent was its lifting body shape.
By flying a lilting trajectory, it was possible to significantly
decrease the amount of gravity losses, thereby
improving vehicle performance and payload capability.
Yet increasing the amount of lift during ascent generally
required flight at higher angles-of-attack and resulted
in greater stress on the vehicle structure. Accordingly,
the nominal trajectory was constrained to keep the parameter
q-_ below a 1500 psf-deg structural design limit
to ensure that the aerodynamic loads did not exceed
the structural capability of the vehicle. The effect of this
trajectory constraint on vehicle performance is shown
in Fig. 3. There was a substantial benefit associated with
using lift during ascent since flying a non-lifting trajectory
resulted in a payload penalty of over 1000 lbs compared
to the nominal case."
http://ntrs.nasa.gov/archive/nasa/casi. ... 025539.pdf

Then the gravity losses could be further reduced by flying a lifting
trajectory, which would also increase the payload capability by a small

The trajectory I'll use to illustrate this will first be straight-line at an
angle up to some high altitude that still allows aerodynamic lift to operate.
At the end of this portion the vehicle will have some horizontal and vertical
component to its velocity. We'll have the vertical component be sufficient to
allow the vehicle to reach 100 km, altitude. The usual way to estimate this
vertical velocity is by using the relation between kinetic energy and
potential energy. It gives the speed of v = sqrt(2gh) to reach an altitude of
h meters. At 100,000 m, v is 1,400 m/s.
Now to have orbital velocity you need 7,800 m/s tangential, i.e., horizontal
velocity. If you were able to fly at a straight-line at a constant angle to
reach 7,800 m/s horizontal velocity and 1,400 m/s vertical velocity and such
that the air drag was kept at the usual low 100 to 150 m/s then you would only
need sqrt(7800^2 + 1400^2) = 7,925 m/s additional delta-v to reach orbit. Then
the total delta-v to orbit might only be in 8,100 m/s range. Note this is
significantly less than the 9,200 m/s delta-v typically needed for orbit,
including gravity and air drag.
The problem is with usual rocket propulsion to orbit not using lift the thrust
vector has to be more or less along the center-line of the rocket otherwise
the rocket would tumble. You can gimbal the engines only for a short time to
change the rocket's attitude but the engines have to be then re-directed along
the center line. However, the center line has to be more or less pointing into
the airstream, i.e., pointing in the same direction as the velocity vector, to
reduce aerodynamic stress and drag on the vehicle. But the rocket thrust
having to counter act gravity means a large portion of the thrust has to be in
the vertical component which means the thrust vector has to be nearly vertical
at least for the early part of the trip when the gross mass is high. Then the
thrust vector couldn't be along the center line of a nearly horizontally
traveling rocket at least during the early part of the trip.
However, using lift you are able to get this large upwards vertical component
for the force on the rocket to allow it to travel along this straight-line. A
problem now though is that at an altitude short of that of space, the air
density will not be enough for aerodynamic lift. Therefore we will use lift
for the first portion of the trajectory, traveling in a straight-line at an
angle. Then after that, with sufficient vertical velocity component attained
to coast to 100 km altitude, we will supply only horizontal thrust during the
second portion to reach the 7,800 m/s horizontal velocity component required
for orbital velocity.

IV.b) How much fuel could we save using a lifting straight-line portion of the
trajectory? I'll give an example calculation that illustrates the fuel savings
from using aerodynamic lift during ascent. First note that just as for
aircraft fuel savings are best at a high L/D ratio. However, the hypersonic
lift /drag ratio of the X-33/VentureStar is rather poor, only around 1.2,
barely better than the space shuttle:

X-33 Hypersonic Aerodynamic Characteristics.
Kelly J, Murphy, Robert J, Nowak, Richard A, Thompson, Brian R, Hollis
NASA Langley Research Center
Ramadas K. Prabhu
Lockheed Martin Engineering &Sciences Company
http://ntrs.nasa.gov/archive/nasa/casi. ... 091447.pdf

This explains the low increase in payload, about 1,000 lbs., less than .5% of
the vehicle dry weight, by using a lifting trajectory for the VentureStar.
However, some lifting body designs can have a lift/drag ratio of from 6 to 8
at hypersonic speeds:

Waverider Design.
http://www.aerospaceweb.org/design/wave ... ider.shtml

The L/D is usually optimized for a specific speed range but we can imagine
"morphing" wings that allow a good L/D ratio over a wide speed range. For
instance note on the "Waverider Design" web page the vehicles optimized for
the highest hypersonic speeds have a long, slender shape, compared to those
for the slower hypersonic speeds. Then for an orbital craft we could have
telescoping sides of the vehicle that would be extended when full of fuel at
the slower speeds, and retracted, producing a slimmer vehicle, when most of
the fuel is burned off and the vehicle is flying faster. Note that a good L/D
ratio at the highest hypersonic speeds means the vehicle will experience less
aerothermal heating on return.
Then we can imagine a second generation lifting trajectory vehicle having this
high L/D ratio over a wide speed range. So in the example I'll take the
supersonic/hypersonic L/D ratio as 5, and for lack of a another vehicle I'll
use the reconfigured kerosene-fueled X-33's thrust and weight values.
Here's the calculation for constant L/D at a constant angle θ (theta).
I'll regard the straight-line path as my X-axis and the perpendicular to this
as the Y-axis. Note this means my axes look like they are at an angle to the
usual horizontal and vertical axes, but it makes the calculation easier. Call
the thrust T, and the mass, M. Then the force component along the
straight-line path, our X-axis, is Fx = T - gMsin(θ) - D and the force
component along the Y-axis is Fy = L - gMcos(θ). We'll set L =
gMcos(θ). Then the force along the straight-line is Fx = T -
gMsin(θ) - gMcos(θ)/(L/D). As with the calculation in the horizontal
case, divide this by M to get the acceleration along this line, and integrate
to get the velocity. The result is V(t) = Ve*ln(M0/Mf) -g*tsin(θ) -
g*tcos(θ)/(L/D), with M0 the initial mass, and Mf, the mass at time t, a
la the rocket equation. If you make the angle θ (theta) be shallow, the
g*tsin(θ) term will be smaller
than the usual gravity drag loss of g*t and the (L/D) divisor will make the
cosine term smaller as well.
I'll assume the straight-line path is used for a time when the altitude is
high enough to use the vacuum Ve of 331s*9.8 m/s^2 = 3244 m/s. According to
the Astronautix page, 3 NK-33's would have a total vacuum thrust of 4,914,000
N and for an Isp of 331s, the propellant flow rate would be
4,914,000/(331x9.8) = 1,515 kg/sec. I'll use the formula: V(t) = Ve*ln(M0/Mf)
- g*tsin(θ) - g*tcos(θ)/(L/D) , to calculate the velocity along the
inclined straight-line path. There are a couple of key facts in this formula.
First note that it includes *both* the gravity and air drag. Secondly, note
that though using aerodynamic lift generates additional, large, induced drag,
this is covered by the fact that the L/D ratio includes this induced drag,
since it involves the *total* drag.
I'll take the time along the straight-line path as 100 sec. Then Mf =
328,700kg -100s*(1,515 kg/s) = 177,200 kg. After trying some examples an angle
of 30º provides a good savings over just using the usual non-lifting
trajectory. Then V(t) = 3244*ln(328,700/177,200) - 9.8*100(sin(30º) +
cos(30º)/5) = 1,345 m/s. Then the vertical component of this velocity is Vy =
1,135*sin(30º) = 672.3 m/s and the horizontal, Vx = 1,135*cos(30º) = 1,164.5
To compare this to a usual rocket trajectory I'll calculate how much fuel
would be needed to first make a vertical trip to reach a vertical speed of
672.3 m/s subject to gravity and air drag, and then to apply horizontal thrust
to reach a 1,164.5 m/s horizontal speed.
The air drag for a usual rocket is in the range of 100 m/s to 200 m/s. I'll
take the air drag loss as 100 m/s for this vertical portion. Then the equation
for the velocity along this vertical part including the gravity loss and the
air drag loss would be V(t) = 3244*ln(M0/Mf) - 9.8*t - 100 m/s, where M0
=328,700 kg and Mf = 328,700 - t(1,515). You want to find the t so that this
velocity matches the vertical component in the inclined case of 672.3 m/s.
Plugging in different values of t, gives for t = 85 sec, V(85) = 680 m/s.
Now to find the horizontal velocity burn. Since this is horizontal there is no
gravity loss, and I'll assume this part is at very high altitude so has
negligible air drag loss. Then the velocity formula is V(t) = 3244*ln(M0/Mf).
Note in this case M0 = 328,700 - 85*1,515 = 199,925 kg, which is the total
mass left after you burned off the propellant during the vertical portion, and
so Mf = 199,925 - t*1,515. Trying different values of t gives for t = 40,
V(40) = 1,171.5 m/s.
Then doing it this way results in a total of 125 sec of fuel burn, 25 percent
higher than in the aerodynamic lift case, specifically 25s*1,515 kg/s = 37,875
kg more. Or viewed the other way, the aerodynamic lift case requires 20% less
fuel over this portion of the trip than the usual non-lift trajectory. With a
307,000 kg total fuel load, this corresponds to a 12.3% reduction in the total
fuel that would actually be needed. Or keeping the same fuel load, a factor
1/.877 = 1.14 larger dry mass could be lofted, which could be used for greater
payload. For a reconfigured X-33 dry mass of 21,700 kg, this means 3,038 kg
extra payload. Remember though this is for our imagined new X-33 lifting shape
that is able to keep a high L/D ratio of 5 at hypersonic speed, not for the
current X-33 shape which only has a hypersonic L/D of 1.2.
With the possibility of using morphing lifting body or wings with high
hypersonic L/D ratio allowing a large reduction in fuel requirements to orbit,
this may be something that could be tested by amateurs or by the "new space"
launch companies.

V.) Now for the calculation of the payload the VentureStar could carry using
kerosene/LOX engines. The propellant mass of the VentureStar was 1,929,000
lbs. compared to the X-33's 210,000 lbs., i.e., 9.2 times more. Then its
propellant tank volume would also be 9.2 times higher, and the kerosene/LOX
they could contain would also be 9.2 times higher, or to 9.2*307,000 =
2,824,400 kg.
We saw the VentureStar dry mass was 257,000 lbs, 116,818 kg, with half of this
as just the mass of the LH2/LOX tanks, at 138,000 lbs, 62,727 kg. However,
going to kerosene/LOX propellant means the tanks would only have to be 1/100th
the mass of the propellant so only 28,244 kg. Then the dry mass would be
reduced to 82,335 kg. We need kerosene/LOX engines now. I suggest the RS-84 be
completed and used for the purpose. You would need seven of them to lift the
heavier propellant load. They weigh about the same as the aerospike engines on
the current version of the VentureStar so you wouldn't gain any weight savings
To calculate how much we could lift to orbit I'll take the average Isp of the
RS-84 as 320. Then if we took the payload as 125,000 kg the total liftoff mass
would be 2,824,400 + 82,335 + 125,000 = 3,031,735 kg, and the ending dry mass
would be 207,335 kg, for a mass ratio of 14.6. Then the total delta-v would be
3200ln(14.6) = 8,580 m/s. Adding on the 462 m/s equatorial speed brings this
to 9042 m/s. With the reduction in gravity drag using a lifting trajectory
this would suffice for orbit.

Bob Clark


Nice read. Didn't know about RS-84 being one more victim of the last 8 years, but fits in. It might be a bit late to include it in any 5 year plans, but it's nice to know, that NASA was building new engines, they just got canceled before they could really prove themselves. Please don't compare them to that ugly exploding stick.
Seems that kerosene engines are already being considered for a heavy lifter, using RD-180.

This would be more shopping than a research ;)


And SpaceX licensed the RS-84, supposedly to create single engine F9's and triple engine F9H's for DoD payloads. Apparently they don't like clusters.

As for the X-33, can't we let that dead dog lie?


docm":29u1504y said:
And SpaceX licensed the RS-84, supposedly to create single engine F9's and triple engine F9H's for DoD payloads. Apparently they don't like clusters.
Noticed, but missed, touche.

docm":29u1504y said:
As for the X-33, can't we let that dead dog lie?
No ;) We want it to fly ! (at least the three of us: me, myself and i).
Thing is, it could and it should and i think it will, probably under different name.


MeteorWayne":10ge658z said:
This belongs in Space Business and Technology.

The moderators are free to move it if they wish.

Bob Clark


Okay, so SSTO is a great concept. I regard the old Atlas as 'nearly' SSTO i.e. it is a workable concept. Differences being we want it reusable. Would strapping on a couple of solids for the initial push be out of the question?


The main issue I see with this is replacing the XRS2200 with the RS-84's, the Linear aerospike was designed to be a component of the aerodynamics of the X-33.

Additionally the RS-84's are longer and will require additional framework, bodywork and a gimbal for thrust vectoring, potentially adding weight.

This will shift the centre of gravity of the vehicle.

Finally the X33 had no payload bay.


No abort, no thrust control, heavy to transport, ..., no solids. Fly-back version of EELV used as a strap on booster would work i guess, kind of a Baikal.



Great, someone discussing advanced concepts!

Some minor points - scaling up is not necessarily linear, internal structure in a tank (other than any essential slosh baffles) usually isn't useful. Most current LVs climb more or less vertically until above the sensible atmosphere, then pitch over to a horizontal thrust vector; this is generally most efficient. There are situations where lofting (climbing initially to an altitude above the final orbit) helps, but with liquids one can modulate both thrust and heading, and I'm not aware of any case where thrusting in a direction greatly different fromt he velocity vector is of any benefit, at least until one is above the atmosphere where the rocket can be pointed in any direction.

Winged lift in the atmosphere adds drag losses, except for air-launch from an air-breathing aircraft. For solid fuel rockets like the Pegasus, which have considerable thrust, I'm not sure how much the wings add. For liquids, where engine mass is a factor, wings might help; the X-34 would have provided some verifiable facts regarding winged lift in climb-out.

The effectiveness of the aerospike isn't really known and this would have been the most interesting result from the X-33. Don't be too hard on the X-33; it was a technology demonstrator, not a prototype of an actual orbital launch vehicle. The tank design just wasn't reasonable.

I would not necessarily exclude two-stage concepts; essentially the SpaceShip I is a two-stage approach that is very cost-effective. The essential element is operational cost. If the staging can be accomplished without explosive bolts or complex interfacing, it could be perfectly acceptable. The two logical points to stage are subsonic (for air launch from a conventional aircraft like the White Knight) or at the top of the sensible atmosphere, since this is the last point a booster could glide back to the launch site, and above this point the thrust requirement is much lower.

The Atlas came close to SSTO, but the booster engines were a large fraction of the empty weight; nevertheless there is room to improve engine weight; the RS-68 for example has about the same mass as the SSME but twice the thrust.

Finally, the dual-fuel concept (use RP-1 or methane at launch, switch to LH2 at altitude in the same engine) also deserves further examination.


neutrino78x":26zo0fp6 said:
What about X-33 using compression to store the hydrogen, or physical storage, such as carbon buckyballs (C-60)?


Hydrogen storage by reversible bonding to carbon is space-efficient but not weight-efficient; probably LH2 remains a better choice for rocket propulsion. If you are going to bond the hydrogen to carbon, then of course you might as well burn the carbon, as in normal kerosine fuel.

Again, the SpaceX Merlin is the only new hydrocarbon engine for space launch developed in the US since Apollo. How can we have fallen so far behind in such a basic technology? There is still room for improvement, for example the linear aerospike approach hasn't been considered.


The same reconfiguration of the Lockheed version of the X-33 to dense fuels and engines to transform it into a full orbital vehicle would also work for the other proposed half-scale suborbital demonstrators.
The McDonnell-Douglas version was essentially the DC-X, scaled somewhat larger. See the attached image. I don't know how much the McD-D version of the X-33 would have cost. However, according to this Astronautix page a 1/2-scale version of the full orbital DC-Y had been proposed, but not funded, which would have cost in the range $450 million, compared to the $60 million of the DC-X, in 1990's dollars:


This would have just below suborbital to suborbital performance, but the price would be significantly less than the DC-Y full orbital version of $5 billion:


However, the point is some preliminary calculations show this 1/2-scale DC-X2 should be able to carry enough dense hydrocarbon fuel under such a reconfiguration to reach orbit. So you would be able to get a reusable SSTO prototype at a significantly reduced price than the $5 billion suggested for the full DC-Y vehicle program.

Bob Clark

Figure 5: X-33 Concept Art from McDonnell Douglas (Frassanito, J.,
McDonnell Douglas).

taken from:

Single Stage to Orbit:
A Reliable Transport System or an Unattainable Dream?


Still the development cost of such a DC-X2 would be quite high, in the range
of $450 million (1990's dollars). So I was still thinking about how small we
could make a scaled-up, reconfigured DC-X to achieve orbit to the extent that
one of the "New Space" companies could afford to build it. I noticed that in
the DC-X there was a lot of empty space, at least according to the diagrammatic
image on the Astronautix page:


I estimated that if we actually fully used up the conical internal space with
propellant, with just a small area at the top for payload or no internal
payload bay at all, made it of an all composite construction (remember the
DC-X was not weight optimized since it would not even go suborbital) and if
we used highly densified hydrocarbon/LOX propellant, to near the solid phase,
then we could get quite high velocities from the DC-X, perhaps up to Mach 20.
In that case only a small scale up from the original DC-X dimensions would
allow you to reach full orbital performance. This would be much cheaper than
the DC-X2. I'm thinking it might even doable for less than $100 million in
current dollars.
Then this could be doable by one of the New Space companies, particularly
those with deep pockets such as SpaceX, Scaled Composites, XCor, Blue Origin, etc.
The case of Blue Origin is particularly interesting because several of the
DC-X engineers moved over to work for Blue Origin and the design of its "New
Shepard" suborbital craft has been likened to that of the DC-X. Blue Origin's
head Jeff Bezos has also said his intention is to move to orbital craft:

Blue Origin.

Blue Origin New Shepard.


Credit- NASA


credit: Blue Origin

Blue Origin First Flight.


I saw the cables holding this thing up. Seriously, this looks like something out of 2001 space odyssey. If it works we will be able to get delivery of our Amazon products in space.

Ok, more seriously, Bezos is a serious guy. Good luck to him.


Call me a pragmatist, but when somebody finally puts a SSTO vehicle in orbit, I'll buy in. Until then, they are all paper (or powerpoint) vehicles.


MeteorWayne":2d0v29wm said:
Call me a pragmatist, but when somebody finally puts a SSTO vehicle in orbit, I'll buy in. Until then, they are all paper (or powerpoint) vehicles.

I agree and I’m not sure why the Augustine commission didn’t include one of these “PowerPoint Vehicles” as one of their “PowerPoint Options”? :lol:


As it is, reusable, kerosene fueled two stage launcher is very close to jump out of presentation in the orbit, Falcon 9 launch is more or less soon, and X-37B, which could also fit description, will be probably secretly launched about the same time.

This is what's bothering me:
http://en.wikipedia.org/wiki/Reusable_l ... y_concepts
One way to increase the mass ratio is to reduce the mass of the empty vehicle by using very lightweight structures and high efficiency engines. This tends to push up maintenance costs as component reliability can be impaired, and makes reuse more expensive to achieve. The margins are so small with this approach that there is uncertainty whether such a vehicle would be able to carry any payload into orbit. Also, lightweight implies small vehicles, which in turn implies small payloads, increasing the cost per kilogram of the payload.

The other way to increase the mass ratio is to increase the mass of the propellant by building very large vehicles. The square-cube law means that the surface area of the vehicle structure increases with the square of the size of the vehicle, while the mass of the fuel increases with the cube of the size; but the weight of the structure increases at a rate somewhere between the two, so a larger vehicle needs somewhat less weight per enclosed volume. If you make the vehicle large enough you can attain a very large mass fraction no matter how heavy the structure.
Could a heavy launcher be the best solution for a SSTO ?

Some other thoughts about using dense fuels, particularly boron-kerosene mix, which would improve volumetric Isp :

This is an old, but fun read:
http://www.spacefuture.com/archive/beta ... cept.shtml

Anyway, there's a lot of people busy with SSTO vehicles, many are not just a presentation anymore, just not all for the Earth orbit at the moment.
Just a reminder ;)


Thanks for those links, Earthling.
Notice the DC-X conical shape seems to be derived from that of the BETA vehicle design.
Also notice that the BETA design uses conventional construction materials. Switching over to an all composite design could improve the payload capability significantly:

The Policy Origins of the X-33.
December 22, 1999
Part II
"Lee also explained to Goldin that primary structures, such as aeroshell sections, made of graphite composite materials would reduce overall structural weight by 40 percent compared to the Shuttle structure, without corrosion or fatigue problems. Actual transport aircraft with composite primary structures included the Boeing 777, the Airbus A330 and A340, and the ATR 72. In addition, Lee wrote, NASA would study both tri-propellant rocket engines and an upgraded Space Shuttle Main Engine (SSME) for use on a single-stage-to-orbit vehicle."

Also, the BETA uses hydrogen fuel. However, as I and others have argued using dense hydrocarbon fuel can actually increase your payload to orbit despite having a lower Isp than hydrogen.

I strongly urge anyone interested in this hydrogen vs. dense hydrocarbon question for a SSTO to read John Whitehead's paper in detail:

Single Stage To Orbit Mass Budgets Derived From Propellant Density and Specific Impulse.
John C. Whitehead
32nd AIAA/ASME/SAE/ASEE Joint Propulsion Conference Lake Buena Vista, FL July
1-3, 1996
http://www.osti.gov/bridge/servlets/pur ... 379977.pdf

He argues that hydrocarbon fuels are more likely to result in a SSTO vehicle
despite the lower Isp.
Three extremely important facts that *must* be taken into account decide

1.)The question of tankage weight. The problems with the tanks on the X-33 are
now legendary within the space community. But it is important to note that it
started off at a disadvantage even at the start because the unusual shapes
required the hydrogen tanks to be *twice* as heavy as normal cylindrical tanks
and the oxygen ones to be *four* times as heavy. For an SSTO that is an
extreme weight penalty.
Moreover, as Whitehead notes, simply using hydrogen/LOX involves a weight
penalty(!) Even normal cylindrical tanks for hydrogen/LOX only provide a 35 to
1 propellant mass to tank mass ratio, compared to 100 to 1 for kerosene/LOX
Even when you take into account you need about twice as much propellant with
kerosene than with hydrogen because of the lowered Isp, this still means you
need about 30% greater tankage weight for the hydrogen case than for the
kerosene case. For an SSTO that is a huge weight penalty.
2.)The question of engine weight. Kerosene engines are typically lighter for
the thrust they produce than hydrogen ones. As an example the NK-33 had about
a 136 to 1 vacuum thrust to weight ratio while for the shuttle main engines
it's only about 73 to 1:

Thrust-to-weight ratio.
# 4.2 Jet and Rocket Engines.
http://en.wikipedia.org/wiki/Thrust-to- ... et_Engines

For the aerospike engines on the X-33 it was even worse at only about 40 to 1 (!)
Since the gross weight will be about twice as much for the kerosene fueled
vehicle, requiring twice as much thrust, this is about even for the SSME's
compared to the NK-33. However, for the aerospike engine currently on the
X-33, this is a huge advantage for the NK-33, since even when you take into
account the twice as great thrust required, the aerospike engines require
about 50% greater weight.
3.)The question of gravity loss. Because kerosene engines can produce greater
thrust for their weight than hydrogen ones, they can reach altitude more
quickly and will therefore have reduced gravity losses compared to hydrogen
ones. This can result in significant total delta-v savings, from 200 m/s to 300
m/s for vehicles with a low gravity loss compared to those with higher losses.

Note that all three of these penalties apply even when you take into account
the higher Isp of hydrogen

Bob Clark


EarthlingX":1h4gia23 said:
As it is, reusable, kerosene fueled two stage launcher is very close to jump out of presentation in the orbit, Falcon 9 launch is more or less soon, and X-37B, which could also fit description, will be probably secretly launched about the same time.

This is what's bothering me:
http://en.wikipedia.org/wiki/Reusable_l ... y_concepts
One way to increase the mass ratio is to reduce the mass of the empty vehicle by using very lightweight structures and high efficiency engines. This tends to push up maintenance costs as component reliability can be impaired, and makes reuse more expensive to achieve. The margins are so small with this approach that there is uncertainty whether such a vehicle would be able to carry any payload into orbit. Also, lightweight implies small vehicles, which in turn implies small payloads, increasing the cost per kilogram of the payload.

The other way to increase the mass ratio is to increase the mass of the propellant by building very large vehicles. The square-cube law means that the surface area of the vehicle structure increases with the square of the size of the vehicle, while the mass of the fuel increases with the cube of the size; but the weight of the structure increases at a rate somewhere between the two, so a larger vehicle needs somewhat less weight per enclosed volume. If you make the vehicle large enough you can attain a very large mass fraction no matter how heavy the structure.
Could a heavy launcher be the best solution for a SSTO ?

Thanks for making that point, Earthling. Notice that for all three proposals made to NASA for a reusable launch vehicle (RLV) by McDonnell-Douglas, Rockwell, and Lockheed, their half-scale demonstrators were only suborbital, not orbit capable, eventhough they used the same materials and same shapes, just half-sized.
The full-scale RLV's very likely would have been SSTO capable. They would have been costly however. That's why I'm arguing in favor of switching over to dense hydrocarbon propellants. This would allow now even the half-scale demonstrators to become orbital craft. This would result in a reduction in the cost by perhaps a factor of 5 to 10. Instead of it being in the billions, it would be in the range of a few hundred million.

Bob Clark



One of the well-financed New Space companies could develop a small-payload capable, all composite, dense propellant VTVL SSTO. This might still cost ca. $100 million. So perhaps just like NASA wanted a half-scale suborbital demonstrator first, perhaps the New Space companies could do this as well.
That is, since a slightly larger all-composite, weight optimized, dense propellant DC-X might be orbit-capable, perhaps the New Space companies could do a half-scale version of *this*. This should be capable of high Mach, hypersonic velocities and suborbital flight. We might estimate that since the size would be 1/2-scale, the volume and mass might be 1/8, and the cost might therefore be 1/8 of the dense propellant version of the DC-X so in the range of $12 million. This might be an amount the New Space companies might want to take a chance on.
But it would be really great if even the small New Space companies could also investigate this. I'm thinking of companies for example like Armadillo Aerospace and Masten Space Systems that took part in the Lunar Lander X-prize competition. I've read that the costs of carbon fiber composites are dropping markedly, so much so that soon some passenger cars will be brought to market with carbon composites making up a significant portion of their mass, something that previously was restricted to million dollar race cars.
So some of these smaller companies might be able to make some small test vehicles using all composite construction that would confirm the principle that all composite construction can result in such large mass ratios that it is equivalent to having SSTO performance. For such small test vehicles these would not need to be reusable, so could save weight on landing gear, thermal protection, wings or stored propellant for landing, etc. These would just be proof of principle concept vehicles that would suggest that with proper scaling relationships a larger all composite vehicle should be SSTO and reusable. See the discussion of the scaling relationships of orbital vehicles here:

Reusable launch system.
2 Reusability concepts.
2.1 Single stage.
http://en.wikipedia.org/wiki/Reusable_l ... ngle_stage

These New Space companies might be able to keep the costs for these small-scale demonstrators low by doing something I hadn't previously known was possible: making your own carbon composite structures in house.
After a web search I saw that some amateurs use carbon composites to save weight both for home-built aircraft and model aircraft:

Homebuilt aircraft.
http://en.wikipedia.org/wiki/Homebuilt_ ... #Composite

Carbon Fiber Composites.

Then by making their composite structures in house the New Space companies could reduce their costs significantly at least for these small scale test vehicles.

Bob Clark


Guys, it's a simple equation to see why a reusable SSTO vehicle should be possible.
It has been often noted that the 1960's era Titan II first stage in itself had single-stage-to-orbit performance, though it would have had minimal payload capability:

SSTO Cons.
"A SSTO vehicle needs to lift its entire structure into orbit. To reach orbit with a useful payload, the rocket requires careful and extensive engineering to save weight. This is much harder to design and engineer. A staged rocket greatly reduces the total mass that flies all the way into space; the rocket is continually shedding fuel tanks and engines that are now dead weight.
"Although a SSTO rocket might theoretically be built, margins would be likely to be very thin- even comparatively minor problems may tend to mean that a project to achieve this could fail to achieve the necessary mass-fraction to reach orbit with useful payload.
"Single-stage rockets were once thought to be beyond reach, but advances in materials technology and construction techniques have shown them to be possible. For example, calculations show that the Titan II first stage, launched on its own, would have a 25-to-1 ratio of fuel to vehicle hardware.[1] It has a sufficiently efficient engine to achieve orbit, but without carrying much payload.[2]"
http://en.wikipedia.org/wiki/Single-sta ... #SSTO_Cons

See the section on Titan II first stage fully fueled mass and empty mass here:

"Stage1: 1 x Titan 2-1. Gross Mass: 117,866 kg (259,850 lb). Empty Mass: 6,736 kg (14,850 lb). Motor: 2 x LR87-7. Thrust (vac): 2,172.231 kN (488,337 lbf). Isp: 296 sec. Burn time: 139 sec. Length: 22.28 m (73.09 ft). Diameter: 3.05 m (10.00 ft). Propellants: N2O4/Aerozine-50."

The two LR-87-7 engines used had a mass of 713 kg each, for a total of 1426 kg:

"Engine Model: LR87-7. Manufacturer Name: AJ23-134. Government Designation: LR87-7. Designer: Aerojet. Propellants: N2O4/Aerozine-50. Thrust(vac): 1,086.100 kN (244,165 lbf). Thrust(sl): 946.700 kN (212,827 lbf). Isp: 296 sec. Isp (sea level): 258 sec. Burn time: 139 sec. Mass Engine: 713 kg (1,571 lb). Diameter: 1.53 m (5.00 ft). Length: 3.13 m (10.26 ft). Chambers: 1. Chamber Pressure: 47.00 bar. Area Ratio: 9.00. Oxidizer to Fuel Ratio: 1.90. Thrust to Weight Ratio: 155.33. Country: USA. Status: Study 1961. First Flight: 1962. Last Flight: 2003. Flown: 212."

Most of the remaining empty mass of 5310 kg for the Titan II first stage would be structural mass, the propellant tanks, support structures, etc. This would be primarily aluminum and steel. Since a 1/3 to 1/2 weight saving can be made over aluminum and steel by using carbon composites, an all composite construction could save at least 1,770 kg off the vehicle empty mass.
However, probably we would have to swap out the engine because as described here, the LR-87-7 engine was not throttlable, which would be needed for a SSTO:

Newsgroups: sci.space.tech
From: henry@spsystems.net (Henry Spencer)
Subject: Re: Is Roton Dead?
Date: Tue, 9 Jan 2001 21:11:51 GMT

In article <93eaqs$6a4$1@mulga.cs.mu.OZ.AU>,
David Kinny <dnk@OMIT.cs.mu.oz.au> wrote:
>>...in fact, the central problem with using
>>the Titan II first stage as an SSTO is that it has *too much* thrust to
>>fly an efficient trajectory.
>How exactly does too much thrust prevent flying an efficient trajectory?
>Difficulties in flipping over to horizontal? Or something else?

Basically, in the time it takes to climb clear of the atmosphere, it picks
up too much vertical velocity. This thing was an ICBM, designed to move
out fast... and flying as an SSTO, it hasn't got a hulking great second
stage on top to slow it down. (In fact, a secondary problem of having too
much thrust is the bone-crushing acceleration toward the end, when the
tanks are almost empty.) An SSTO launcher wants to take things a bit
slower, so that it can tip over to horizontal gradually, as it leaves the
atmosphere, and still have most of its fuel left for horizontal acceleration.
You can't just throttle back the engine, first because it wasn't
throttlable :), and second because you need to keep it operating
efficiently, which throttling usually sacrifices to at least some extent.
However, *reducing* the performance of an engine is usually not a
difficult engineering problem!
When failure is not an option, success | Henry Spencer henry@****.net
can get expensive. -- Peter Stibrany | (aka ****@zoo.toronto.edu)

I suggest the NK-33 be used. It was designed for kerosene/LOX but quite likely would also work with the N2O4/Aerozine-50 propellant of the LR-87-7 because the LR-87 engine was variously used with N2O4/Aerozine-50 and kerosene/LOX. Using a single NK-33 would also save 200 kg off the vehicle dry weight:


The question is could we use that approx. 2,000 kg saved weight for landing gear and thermal protection to make the vehicle reusable? Let's take the landed weight as still 6,736 kg where we used the saved weight for landing gear, thermal protection, and payload.
The landing gear for an aerial vehicle is commonly taken as 3% of the landed weight:

Landing gear weight.
http://yarchive.net/space/launchers/lan ... eight.html

So this is 202 kg.
To make a powered vertical landing the common estimate is 10% of the vehicle landed weight has to be used in propellant:

Reusable launch system.
Vertical landing.
http://en.wikipedia.org/wiki/Reusable_l ... al_landing

So 673 kg.

For thermal protection, we'll assume it'll make a ballistic reentry, base first. For this vehicle, the base will only be 3 meters wide, for an area of 7 m^2. Using base first reentry we'll have to cover primarily the base only:

Blue Origin New Shepard.
"A passenger and cargo spacecraft has considerably less need for cross-range."
"As a result, the craft is much "rounder" than the DC-X, optimized for tankage and structural benefits rather than re-entry aerodynamics. It has not been stated if the vehicle is intended to re-enter base-first or nose first, but the former is most likely for a variety of reasons. For one, it reduces heat shield area, and thus weight, covering only the smaller bottom surface rather than the much larger upper portions. The area around the engines would likely require some sort of heat protection anyway, so by using the base as the heat shield the two can be combined. This re-entry attitude also has the advantage of allowing the spacecraft to descend all the way from orbit to touchdown in a base-first orientation, which would seem to offer some safety benefits as well as reducing aero-loading issues."

We'll use the high temperature resistant but low maintenance metallic shingles developed for the X-33:

http://reference.kfupm.edu.sa/content/r ... 117853.pdf

The areal density of this is in the range of 10 to 15 kg/m^2. This will then require 70 to 105 kg to cover the base only.
Then the total mass for landing and thermal protection up to 980 kg, and about 1,000 kg could go to payload. This would be only 0.8% of the gross mass but would be for a reusable SSTO vehicle.
It might be possible to improve this payload fraction by using kerosene/LOX instead of the N2O4/Aerozine-50 propellant. This would result in a higher Isp, however the N2O4/Aerozine-50 is denser and so more fuel can be carried.

Bob Clark


It is important to remember that single-stage-to-orbit in itself is not impossible. It was in fact proven to be feasible from the early days of the space program:

SSTO Cons.
"Single-stage rockets were once thought to be beyond reach, but advances in materials technology and construction techniques have shown them to be possible. For example, calculations show that the Titan II first stage, launched on its own, would have a 25-to-1 ratio of fuel to vehicle hardware.[1] It has a sufficiently efficient engine to achieve orbit, but without carrying much payload.[2]"
http://en.wikipedia.org/wiki/Single-sta ... #SSTO_Cons

Such a vehicle of course while carrying minimal payload would also not be reusable. The question is could you replicate this performance using lightweight materials so this weight savings could go to reentry and return systems and could this be done economically?
I already gave the argument that the weight savings possible from composite construction makes such a reusable SSTO possible. The reason why I say it is now economically feasible is because lightweight carbon composite construction is now being planned for some passenger cars. Consider the price for carbon composites in the early 90's:

G.M. to Show a High-Mileage Experimental Car
Published: Monday, December 30, 1991
"At the North American International Auto Show in Detroit next week, G.M. will show its Ultralite, which the company says can produce 100-mile-a-gallon fuel efficiency at 50-mile-an-hour highway speeds.
"That efficiency is possible, G.M. said, because the car weighs only 1,400 pounds. (A Chevrolet Corsica, which is approximately the same size as the Ultralite, weighs about twice as much.) Scaled Composites Inc. of Mojave, Calif., built the Ultralite body for G.M.
"Although many race cars are made of carbon fiber, which is quite sturdy, the material is enormously expensive compared with steel or aluminum. But G.M. said it had received a patent for a process that sharply reduces the cost of carbon fiber, which currently is about $40 a pound, compared with about 35 cents a pound for steel."
http://www.nytimes.com/1991/12/30/busin ... l-car.html

So the price then was about 100 times greater than steel. You wouldn't see many all-composite-construction rockets at those prices even if even then it would have made a reusable SSTO possible. Now look at the price given in this article from the year 2000:

Carbon-Fiber Composites for Cars.
"To meet the ultimate PNGV mileage goal, one potentially enabling technology is to use carbon-fiber composites, which form the structure of U.S. fighter jets. Carbon-fiber composites weigh about one-fifth as much as steel, but can be comparable or better in terms of stiffness and strength, depending on fiber grade and orientation. These composites do not rust or corrode like steel or aluminum. Perhaps most important, they could reduce vehicle weight by as much as 60%, significantly increasing vehicle fuel economy.
"The problem is that carbon-fiber composites cost at least 20 times as much as steel, and the automobile industry is not interested in using them until the price of carbon fiber drops from $8 to $5 (and preferably $3) a pound. Production of carbon fibers is too expensive and slow."

Now this British company claims their patented process allows composite construction both for the chassis frame and the body panels at low cost for a passenger car to be introduced next year:

Axon announces affordable, 100mpg, carbon-composite passenger car.
"Axon has gone simply for an uncomplicated 500cc engine in a low-weight body, which replaces the traditional heavy steel or aluminium frame with recycled carbon fibre composites - as strong as steel but only around 40% as heavy. Extensive use of carbon materials through Axon’s cars makes a massive impact on the power-to-weight ratio, meaning they can get acceptable overall performance using a much smaller, lighter and more frugal engine.
"The lightness and strength of carbon fibre have been well-known for decades - it’s been cost that’s prevented this wonder-material from popping up all over the automotive world, restricting it to top-end specials and aftermarket goodies. But it’s here that Axon claim to have made a breakthrough."
http://www.transport20.com/uncategorize ... enger-car/

Because of the rate at which the costs of carbon composite production is decreasing, I argue the production cost for a reusable SSTO using carbon composite construction, because the lighter weight in materials required, will soon be comparable to that of an expendable rocket using standard, heavy construction materials. And it is already now economically feasible due to lower per use costs of a reusable vehicle.

It is also extremely important to keep in mind that such a reduction in structural mass for a rocket would result in a comparable reduction in engine mass. This is important because the engine mass is the second greatest component for the dry mass of the rocket after the structural mass.
The reason this engine mass reduction occurs is exactly analogous to why it occurs when replacing the structural mass of cars with lighter materials:

Carbon Fibre Reinforced Composite Car.
Primary author: Andrew Mills
Source: Materials World, Vol 10, no. 9 pp. 20-22, September 2002.
"In the area of vehicle design, body weight is the most important target for improvement, as a reduction in the weight of a vehicle’s body means that a smaller engine, and a lighter drive train and assembly can be used. This ‘benign spiral’ leads to further mass reductions, so much so that various studies have indicated a potential for savings of up to 65% by using carbon fibre composites instead of steel wherever possible."
http://www.azom.com/Details.asp?Article ... Background

Bob Clark



LORAL Space Systems, the leading communications satellite builder,
had a design for a single-stage-to-orbit though expendable launcher.
They expected to use all-composite cryogenic tanks on these launchers
to save weight. Their idea was that the high cost of launch is from
trying to assure high reliability. However, their launchers were to be
designed to be used for payloads such as replacing consumables on the
ISS, launching propellants to orbital depots, etc.
They were able to conclude based on study of prior launchers that
high reliable launchers cost more and correspondingly lower reliable
ones cost less. They therefore specifically aimed for a rather low
reliability rate of about 66% to get low cost. They figured this would
be allowable for low cost items such fuel and consumables.
Still, it is interesting that their low cost design was specifically
based on a SSTO, composite-tank rocket:

Aquarius: Low-Cost Low Reliability Consumables Launcher.
Enabling Technology includes large, lightweight liner-less composite tanks.
http://homepage.mac.com/fcrossman/NorCa ... 012204.pdf

"Proposed expendable, water launch, single-stage-to-orbit, liquid
oxygen/hydrogen, low-cost launch vehicle designed to carry small bulk
payloads to low earth orbit. A unique attribute was that low
reliability was accepted in order to achieve low cost."

Bob Clark
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