A kerosene-fueled X-33 as a single stage to orbit vehicle.
Robert Clark
I.)Introduction.
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/purl/379977-2LwFyZ/webviewable/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 kerosene/LOX.
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."
http://www.space-access.org/updates/sau91.html
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.
http://www.xs4all.nl/~carlkop/x33.html
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/OtherPRINT/Lithographs/X33.Advanced.Technology.Demonstrator.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 forum:
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 velocity.
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_mechanics/plates/calculators/SSSS_PUniform.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 strength*.
A possible way tensile strength could be provided would be to use the
Poisson's ratio of the honeycombed cells:
Poisson's ratio.
http://en.wikipedia.org/wiki/Poisson%27s_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.
http://silver.neep.wisc.edu/~lakes/PoissonChiral.html
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/RT2001/5000/5870tomsik.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:
RS-84.
http://www.astronautix.com/engines/rs84.htm
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/membersonly/october98/features/launch/launch.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:
NK-33.
http://www.astronautix.com/engines/nk33.htm
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/2008/01/10/drag-loss-in-ascent-gain-in-descent-and-what-it-means-for-scalability/
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.ntrs.nasa.gov/20000031364_2000025539.pdf
Then the gravity losses could be further reduced by flying a lifting
trajectory, which would also increase the payload capability by a
small percentage.
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 the 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:
AIAA-99-4162
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.ntrs.nasa.gov/20040087108_2004091447.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/waverider/waverider.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 also 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
for the usual rocket equation, 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 m/s.
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 fomula 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, thus 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 here.
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
Sam Wormley
Pat Flannery
> 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.
You are missing a key point here; Lockheed chose LOX/LH2 for two reasons:
1.) Its superior isp.
2.) The large size of the propellant tankage for it.
VentureStar was to rely on the large volume of the empty propellant
tankage to make the vehicle very light for its size on reentry to
decelerate the spacecraft quickly and reduce the heat loads on the TPS.
Shift to kerosene/LOX and you are going to need a far heavier TPS, and
combined with the lower isp, that is going to make SSTO very difficult
to do, especially with a worthwhile payload.
Pat
Uncle Al
>
> Table of Contents.
> I.)Introduction.
> 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.
No, you hopeless ignorant ineducable idiot,
exo-tetrahydrodicyclopentadiene [exo-tricyclo[5.2.1.0^(2,6)]decane)
for its liquid high density and awesome enthalpy of combustion.
> 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.
[snip 500 lines]
Thank you for verifying that you are a technological dunce even about
~C_10 hydrocarbon fuels. Bulk synthesis of cyclobutyl and especially
cyclopropyl polycyclic hydrocarbons affords kerosene liquids even
denser and of higher delivered combustion enthalpies than JP-10 for
their ring strain.
Uncle Al, being of delicate aesthetics, proposes adding 1%
8-exo-formyl-2,6-exo-tricyclo[5.2.1.0^(2,6)]decane to JP-10 for its
having a fresh-green note scent and high stability.
Why isn't NASA eviscerated for its obscene carbon footprint?
--
Uncle Al
http://www.mazepath.com/uncleal/
(Toxic URL! Unsafe for children and most mammals)
http://www.mazepath.com/uncleal/qz4.htm
Sylvia Else
> Single Stage to Orbit really limits payload "weight".
Why does that matter? The point of an SSTO is to get down the cost to
orbit per payload kg. An SSTO is likely to mass more than a disposable
multi-stage for a given payload, but that is not in itself a source of
concern.
Sylvia.
Sam Wormley
Sylvia Else
And...?
Sylvia.
Sam Wormley
If it can be done, great! No Earth-launched SSTO launch vehicles have
ever been constructed.
Sylvia Else
They never will be if we use that as an argument against them.
Sylvia.
Sam Wormley
What usually comes first is the desire to put a heavy payload into orbit
around the earth, sun, or another planet. What rocket configuration can
get the job done reliably? That's usually the question, as opposed to
what rocket can get the job done for the least cost per payload kg.
Sylvia Else
Depends which market you're in. Space tourism, for example, needs low
cost per kg, but not particularly large payloads.
Sylvia.
Sam Wormley
I would think that space tourism, would require some margin--some extra
safety.
"It is extremely difficult to design a structure which is strong, safe,
very light, and economical to build. Designers often liken the task
to designing and building an egg shell. The problem originally seemed
insuperable, and drove all early designers to multistage rockets".
"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".
Sylvia Else
> 'it can't be done' thing right off the table.
Yes, even a technology demonstrator with an uneconomic payload would be
a useful starting point. Isn't that the kind of thing NASA is meant to do?
Sylvia.
Sam Wormley
> Sam Wormley <swor...@mchsi.com> wrote:
>
> :Sylvia Else wrote:
> :>
> :> Depends which market you're in. Space tourism, for example, needs low
> :
> : I would think that space tourism, would require some margin--some extra
> : safety.
> :
>
> Why? You need what you need. Why would tourists require 'extra'?
>
>
You lose a human in a rocket, people will want their money back.
Sylvia Else
> "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".
>
I wonder how far aviation would have got if the first aircraft had been
required to be economically viable.
Sylvia.
Sylvia Else
Only if they haven't read the smallprint.
Sylvia.
Sam Wormley
>
> You lose a billion dollar satellite in a rocket and they're going to
> want their money back, too. That's what insurance and liability
> waivers are for.
>
Agreed--And they will probably use more reliable rockets too.
Sam Wormley
> Sam Wormley <swor...@mchsi.com> wrote:
>
> :Fred J. McCall wrote:
> :> Sam Wormley <swor...@mchsi.com> wrote:
> :>
>
> So we're to "all rockets will be reliable". So why do you think
> "space tourism, would require some margin--some extra safety", again?
>
Let me clarify.
1. SSTO launch vehicles have never been used to put things in orbit.
2. SSTOs are hard to design.
3. I'm guessing the are not the best candidate for space tourism.
Sam Wormley
>
> You mean "let me move the goal posts", don't you? So it has nothing
> to with "extra safety" at all.
>
> Now that that's out of the way, let's look at the mission requirements
> for a vehicle intended for 'space tourism':
>
> 1) No more likely to blow up than any other vehicle.
>
> 2) Fast cycle times (space tourism needs to fly frequently to get
> costs down to the point where it's practical).
>
> 3) Low support costs (we're looking at an airliner model of costs
> here, not what is currently happening with present launch vehicles).
>
> 4) Cargo fraction can be low, since the bulk of your 'cargo' is meat
> and seats. You're not staying up long enough to need to worry too
> much about consumable supplies other than air and a minimum of fluids.
>
> So what does that lead us to?
>
> 1) Liquid fuel. Failure rates are similar to solids, but
> instantaneous catastrophic failures are much rarer. Plus the
> vibration environment is much better.
>
> 2) Reusable. If you have to buy new hardware for every shot, there's
> a very real limit to how low you can drive costs and you'll never get
> anywhere near airline cost models.
>
> 3) SSTO or air launched. 'Winged flyback booster' falls in this
> category, although the technology there is more difficult than
> launching from a manned aircraft. If we want fast cycle times, we
> can't just be dropping pieces randomly. Plus having to retrieve them
> rather than having them come home on their own increases cycle costs.
>
> 4) Don't push the technology too hard. Doing that gets you better
> performance, but it also leads to higher refurbishment costs between
> flights as pushing limits harder means closer tolerances and running
> things closer to failure points.
>
> Conclusion: What we're looking for for space tourism (or a personnel
> transport vehicle) is a reusable liquid fueled SSTO (or air-launched
> vehicle) that can fly back to the launch facility and land, then take
> off for another flight with minimal processing.
>
> Seems obvious to me. Also to pretty much everyone else talking about
> doing space tourism. If you think you know better, perhaps you should
> get some backers and start a company?
>
Go for it!
Sylvia Else
> 3) SSTO or air launched. 'Winged flyback booster' falls in this
> category, although the technology there is more difficult than
> launching from a manned aircraft. If we want fast cycle times, we
> can't just be dropping pieces randomly. Plus having to retrieve them
> rather than having them come home on their own increases cycle costs.
Also, not dropping bits off gives you much more flexibility about where
you launch from, particularly if you get the chance of raining burning
debris over inhabited areas down to somewhere around airliner numbers
(whether or not the rest of the mission is that safe).
Sylvia.
Robert Clark
The very key aspect of this proposal is that the tanks remain the
*same* size, but at a *lighter* weight. In fact the intent was to keep
the same shape of the X-33 and just switch out the propellant tanks
and engines. So in fact the vehicle becomes lighter for its volume
with hydrocarbon fuels.
When you consider the other benefits of hydrocarbon fuels over
hydrogen, the higher Isp of hydrogen/LOX propellant becomes less of an
advantage.
In fact, kerosene is not necessarily the best hydrocarbon to use,
which I'll discuss in a following post.
Bob Clark
Robert Clark
Thanks for that link. It also neatly discuses the reasons why a dense
hydrocarbon fuel can be superior to hydrogen for a SSTO, particularly
for the savings in tank weight and gravity losses:
Dense versus hydrogen fuels.
http://en.wikipedia.org/wiki/Single-stage-to-orbit#Dense_versus_hydrogen_fuels
In regards to the payload to orbit, in my suggested reconfigured
kerosene-fueled VentureStar it would exceed the payload of the Saturn
V, and nearly match that of the Ares V.
And with even more efficient hydrocarbon fuels than kerosene it would
exceed even that of the Ares V.
Bob Clark
Pat Flannery
>
> I wonder how far aviation would have got if the first aircraft had been
> required to be economically viable.
When designing the Flyer, it was the intention of the Wright brothers to
make a lot of money by selling them to the US and other governments for
use as military reconnaissance devices.
That's the reason they took out so many patents on it, as it was their
intention to lock any competitors out of the profits that could be
accrued that way.
They were doing pretty good at that tactic until Glenn Curtiss developed
aerilons to replace the Wright wing-warping system for roll control, and
that was not judged as a infringement of the Wright patents.
Pat
Pat Flannery
> The very key aspect of this proposal is that the tanks remain the
> *same* size, but at a *lighter* weight. In fact the intent was to keep
> the same shape of the X-33 and just switch out the propellant tanks
> and engines. So in fact the vehicle becomes lighter for its volume
> with hydrocarbon fuels.
The mixture rate by volume is way different for LOX/LH2 and
Kerosene/LOX, so the tanks well have to be changed in proportional size
to each other.
Also, due to the far higher density of the Kerosene versus LH2, during
acceleration a tank strong enough to carry LH2 will rupture under the
higher weight of Kerosene.
The whole point of the composite tanks on VentureStar was to get their
weight down to lower than ones made out of aluminum, so I'm keen to see
what you are going to make them out of that is lighter than composite
materials.
> When you consider the other benefits of hydrocarbon fuels over
> hydrogen, the higher Isp of hydrogen/LOX propellant becomes less of an
> advantage.
> In fact, kerosene is not necessarily the best hydrocarbon to use,
> which I'll discuss in a following post.
I'm breathless with anticipation.
Pat
Jeff Findley
"Sam Wormley" <swor...@mchsi.com> wrote in message
news:B_fHm.110690$5n1.67859@attbi_s21...
> Single Stage to Orbit really limits payload "weight".
I think you meant to say that SSTO would limit the mass fraction of the
vehicle which is payload. I'd like to note that is not necessarily a bad
thing, despite the belief of traditional aerospace engineers to the
contrary. What you *really* want to optimize is cost per kg to orbit, not
the payload mass fraction of your vehicle. Since the current launch market
is rather small, there hasn't been much effort made in this area by the
traditional launch providers. But I'd also like to note that some start-ups
have been making some progress in this area, even though they're not
actively pursuing SSTO vehicles.
The development costs for a reusable SSTO would be high and the reality is
that the current launch market just isn't big enough to justify the
investment.
Jeff
--
"Take heart amid the deepening gloom
that your dog is finally getting enough cheese" - Deteriorata - National
Lampoon
Sylvia Else
> The military
> wants a missile defense base on the south pole of the Moon, which
> is the only place the Earth can be continually observed.
Seems a long way out for a defense base that can in any case only see
just under 50% of the Earth at a time, and has most of Earth out of view
for nearly 13 consecutive hours in each 25.
Sylvia.
Robert Clark
vehicle using NK-33 engines. The NK-33 was a Russian 1960's era engine
so I thought they would have to be taken out of mothballs for the
purpose. But I recently found that Aerojet is working with the NK-33
engines to be used on Orbital Sciences’ Taurus 2 launcher:
08/31/09 10:15 AM ET
Aerojet Looking to Restart Production of NK-33 Engine.
By Amy Klamper
http://www.spacenews.com/launch/aerojet-looking-restart-production-nk-33-engine.html
August 19, 2009
Russian Mail-Order Ride.
http://blogs.airspacemag.com/daily-planet/2009/08/19/russian-mail-order-ride/
Aerojet has already purchased several of the engines, and is debating
whether to start it's own production lines or to use Russian
production for future purchases of the engine. Then the Air Force or
NASA could use these to make a reusable, single-stage-to-orbit vehicle
and near term, at least for a prototype vehicle. There was some debate
on the Augustine commission if NASA and the U.S. should use Russian
engines for a significant portion of their launches, but this
complaint might be ameliorated in regards to the NK-33 if production
lines were started in the U.S.
A question would be of the payload it could carry. The preliminary
calculations I made suggested it might just make orbit, so likely it
would have low payload capability. Some possibilities to increase the
payload might be to densify the kerosene propellant by subcooling to
near LOX temperatures or to use more energetic hydrocarbon
propellants. The densification would allow it carry more propellant.
Some possible energetic hydrocarbon propellants are suggested here:
Alternate Propellants for SSTO Launchers.
Dr. Bruce Dunn
http://www.dunnspace.com/alternate_ssto_propellants.htm
Another possible method to increase payload would to use a version of
the NK-33 with an aerospike nozzle. This would allow it to have higher
Isp at sea level.
It should also be possible to use the hydrocarbon fueled X-33 as the
a reusable first stage booster. The Air Force is investigating such
boosters as a means of cutting costs to space. Since the reconfigured
X-33 would be able to reach orbit at 21,700 kg dry mass, it could be
able to lift in the range of a few thousand kg's payload as the first
stage of two stage-to-orbit-system.
Bob Clark
On Nov 1, 8:20 am, Robert Clark <rgregorycl...@yahoo.com> wrote:
> Table of Contents.
> I.)Introduction.
> 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, 1996http://www.osti.gov/bridge/servlets/purl/379977-2LwFyZ/webviewable/37...
>
> 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 kerosene/LOX.
> 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."http://www.space-access.org/updates/sau91.html
>
> 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.http://www.xs4all.nl/~carlkop/x33.html
>
> 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/OtherPRINT/Lithographs/X33.Advan...
>
> 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 forum:
>
> http://www.bautforum.com/space-exploration/86728-passenger-market-sub...
>
> 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 velocity.
> 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_mechanics/plates/calculators/SSS...
>
> 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 strength*.
> A possible way tensile strength could be provided would be to use the
> Poisson's ratio of the honeycombed cells:
>
> Poisson's ratio.http://en.wikipedia.org/wiki/Poisson%27s_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.http://silver.neep.wisc.edu/~lakes/PoissonChiral.html
>
> 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. ...
>
> read more »
Robert Clark
> The proposal was to transform the X-33 into a reusable, orbital
> vehicle using NK-33 engines. The NK-33 was a Russian 1960's era engine
> so I thought they would have to be taken out of mothballs for the
> purpose. But I recently found that Aerojet is working with the NK-33
> engines to be used on Orbital Sciences’ Taurus 2 launcher:
>
> 08/31/09 10:15 AM ET
> Aerojet Looking to Restart Production of NK-33 Engine.
>
> August 19, 2009
> Russian Mail-Order Ride.http://blogs.airspacemag.com/daily-planet/2009/08/19/russian-mail-ord...
>
> Aerojet has already purchased several of the engines, and is debating
> whether to start it's own production lines or to use Russian
> production for future purchases of the engine. Then the Air Force or
> NASA could use these to make a reusable, single-stage-to-orbit vehicle
> and near term, at least for a prototype vehicle. There was some debate
> on the Augustine commission if NASA and the U.S. should use Russian
> engines for a significant portion of their launches, but this
> complaint might be ameliorated in regards to the NK-33 if production
> lines were started in the U.S.
> A question would be of the payload it could carry. The preliminary
> calculations I made suggested it might just make orbit, so likely it
> would have low payload capability. Some possibilities to increase the
> payload might be to densify the kerosene propellant by subcooling to
> near LOX temperatures or to use more energetic hydrocarbon
> propellants. The densification would allow it carry more propellant.
> Some possible energetic hydrocarbon propellants are suggested here:
>
> Alternate Propellants for SSTO Launchers.
>
> Another possible method to increase payload would to use a version of
> the NK-33 with an aerospike nozzle. This would allow it to have higher
> Isp at sea level.
>
> It should also be possible to use the hydrocarbon fueled X-33 as the
> a reusable first stage booster. The Air Force is investigating such
> boosters as a means of cutting costs to space. Since the reconfigured
> X-33 would be able to reach orbit at 21,700 kg dry mass, it could be
> able to lift in the range of a few thousand kg's payload as the first
> stage of two stage-to-orbit-system.
>
Aerojet claims their version of the NK-33 is "fully reusable":
Space Lift Propulsion.
http://www.aerojet.com/capabilities/spacelift.php
Anyone have any idea how many reuses they mean by that?
Bob Clark
Pat Flannery
> Aerojet claims their version of the NK-33 is "fully reusable":
>
> Space Lift Propulsion.
> http://www.aerojet.com/capabilities/spacelift.php
>
> Anyone have any idea how many reuses they mean by that?
NK-33 engines were designed for a total burn time of 1,200 seconds (20
minutes): http://www.astronautix.com/engines/nk33.htm
Pat
Robert Clark
> vehicle using NK-33 engines. The NK-33 was a Russian 1960's era engine
> so I thought they would have to be taken out of mothballs for the
> purpose. But I recently found that Aerojet is working with the NK-33
> engines to be used on Orbital Sciences’ Taurus 2 launcher:
>
> 08/31/09 10:15 AM ET
> Aerojet Looking to Restart Production of NK-33 Engine.
>
> August 19, 2009
> Russian Mail-Order Ride.http://blogs.airspacemag.com/daily-planet/2009/08/19/russian-mail-ord...
>
> Aerojet has already purchased several of the engines, and is debating
> whether to start it's own production lines or to use Russian
> production for future purchases of the engine. Then the Air Force or
> NASA could use these to make a reusable, single-stage-to-orbit vehicle
> and near term, at least for a prototype vehicle. There was some debate
> on the Augustine commission if NASA and the U.S. should use Russian
> engines for a significant portion of their launches, but this
> complaint might be ameliorated in regards to the NK-33 if production
> lines were started in the U.S.
> A question would be of the payload it could carry. The preliminary
> calculations I made suggested it might just make orbit, so likely it
> would have low payload capability. Some possibilities to increase the
> payload might be to densify the kerosene propellant by subcooling to
> near LOX temperatures or to use more energetic hydrocarbon
> propellants. The densification would allow it carry more propellant.
> Some possible energetic hydrocarbon propellants are suggested here:
>
> Alternate Propellants for SSTO Launchers.
>
> Another possible method to increase payload would to use a version of
> the NK-33 with an aerospike nozzle. This would allow it to have higher
> Isp at sea level.
>
> It should also be possible to use the hydrocarbon fueled X-33 as the
> a reusable first stage booster. The Air Force is investigating such
> boosters as a means of cutting costs to space. Since the reconfigured
> X-33 would be able to reach orbit at 21,700 kg dry mass, it could be
> able to lift in the range of a few thousand kg's payload as the first
> stage of two stage-to-orbit-system.
>
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 linked 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:
DC-X2.
http://astronautix.com/lvs/dcx2.htm
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:
DC-Y.
http://astronautix.com/lvs/dcy.htm
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).
http://vorlon.case.edu/~jam64/images/SSTO/SSTO_Figure_5.jpg
taken from:
Single Stage to Orbit:
A Reliable Transport System or an Unattainable Dream?
http://vorlon.case.edu/~jam64/work/ssto.htm
Robert Clark
> DC-X2.http://astronautix.com/lvs/dcx2.htm
>
> 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:
>
> DC-Y.http://astronautix.com/lvs/dcy.htm
>
> 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).http://vorlon.case.edu/~jam64/images/SSTO/SSTO_Figure_5.jpg
>
> taken from:
>
> Single Stage to Orbit:
> A Reliable Transport System or an Unattainable Dream?http://vorlon.case.edu/~jam64/work/ssto.htm
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 on
the DC-X there was a lot of empty space, at least according to the
diagrammatic
image on the Astronautix page:
DC-X.
http://www.astronautix.com/lvs/dcx.htm
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.
http://en.wikipedia.org/wiki/Blue_Origin
Blue Origin New Shepard.
http://en.wikipedia.org/wiki/Blue_Origin_New_Shepard
Bob Clark
Robert Clark
> ...
>
> 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.http://en.wikipedia.org/wiki/Blue_Origin
>
> Blue Origin New Shepard.http://en.wikipedia.org/wiki/Blue_Origin_New_Shepard
>
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_launch_system#Single_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_aircraft#Composite
Carbon Fiber Composites.
http://winshiprc.tripod.com/carbon_fiber_composites.htm
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
Robert Clark
> ...
> 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 linked 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.,
>
> taken from:
>
> Single Stage to Orbit:
> A Reliable Transport System or an Unattainable Dream?http://vorlon.case.edu/~jam64/work/ssto.htm
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-stage-to-orbit#SSTO_Cons
See the section on Titan II first stage fully fueled mass and empty
mass here:
Titan.
"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."
http://www.astronautix.com/lvs/titan.htm
The two LR-87-7 engines used had a mass of 713 kg each, for a total of
1426 kg:
LR87.
"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."
http://www.astronautix.com/engines/lr87.htm
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: he...@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 <d...@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)
http://yarchive.net/space/launchers/ssto.html
--------------------------------------------------------------
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:
NK-33.
http://www.astronautix.com/engines/nk33.htm
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/landing_gear_weight.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_launch_system#Vertical_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."
http://en.wikipedia.org/wiki/Blue_Origin_New_Shepard
We'll use the high temperature resistant but low maintenance metallic
shingles developed for the X-33:
REUSABLE METALLIC THERMAL PROTECTION SYSTEMS DEVELOPMENT,
http://reference.kfupm.edu.sa/content/r/e/reusable_metallic_thermal_protection_sys_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 is 980 kg, and
about 1,000 kg could go to payload. This would be only 0.8% of the
gross mass but would be 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
Robert Clark
not impossible. It was in fact proven to be feasible from the early
days of the space program:
Single-stage-to-orbit.
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-stage-to-orbit#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
By DORON P. LEVIN,
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/business/gm-to-show-a-high-mileage-experimental-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."
http://www.ornl.gov/info/ornlreview/v33_3_00/carbon.htm
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/uncategorized/axon-announces-affordable-100mpg-carbon-composite-passenger-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?ArticleID=1662#_Background
Bob Clark
Axontex chasis.
http://www.transport20.com/gallery/albums/Axon/axontex_chassis.jpg
Robert Clark
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/NorCalSAMPE/Comp_WS_papers/Turner_012204.pdf
Aquarius.
"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."
http://www.astronautix.com/lvs/aquarius.htm
Bob Clark
Robert 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
>
> Aquarius.
> "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."http://www.astronautix.com/lvs/aquarius.htm
>
> Bob Clark
Nice list of launch vehicle designs, including some SSTO's going back
to the 60's:
Space Future - Vehicle Designs.
http://www.spacefuture.com/vehicles/designs.shtml
Here's a review of SSTO concepts proposed over the years:
History of the Phoenix VTOL SSTO and Recent Developments in Single-
Stage Launch Systems.
Gary C Hudson
http://www.spacefuture.com/archive/history_of_the_phoenix_vtol_ssto_and_recent_developments_in_single_stage_launch_systems.shtml
And this article argues that SSTO performance has long been possible
for expendables, and that a reusable one is possible with modern
materials:
Launch Vehicle Design.
"Contrary to what many people who make expendable rockets will tell
you, it isn't difficult to design a "single stage to orbit" ( SSTO)
rocket. In fact it's very easy - it can be done with rocket engines
and propellant tanks designed, manufactured and operated 20 years ago!
It's important to know this, because a lot of people will try to tell
you otherwise.
"A Thought Experiment
"This very idea was written up by Gary Hudson in "A Single-Stage-to-
Orbit thought experiment".
"If you attach 6 SSMEs (Space Shuttle Main Engines) directly to a
Space Shuttle External Tank ( ET), you could launch 30 tons payload to
orbit. It wouldn't be an economical way to launch - but it's certainly
possible. But please note: it's only possible taking off vertically;
no-one can build a horizontal take-off SSTO.
"But, of course, if you carry passengers to orbit you'll want to bring
them back - and that's what's tricky: to build a fully reusable SSTO,
not an expendable, one-way ride. "
http://www.spacefuture.com/vehicles/building.shtml
As this article notes, many people don't think a SSTO vehicle is
possible even with expendables. That is why with the rapid drop in the
cost of composite materials I'm arguing that small test vehicles of
all-composite construction should be built to prove the principle of
SSTO at least for expendables. This would be possible and affordable
to do even for the smallest of the New Space companies with in house
construction of the composite materials.
Then when it is seen that SSTO, though not reusable, performance is
possible for an actual working rocket, it will be more believable that
following well-known scaling principles that larger rockets should
allow reusable versions with significant payloads.
Bob Clark
Robert Clark
> 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
>
> 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.
>
Video of SpaceShipTwo assembly, showing the all-composite
construction, including the structural members:
SpaceShipTwo Assembly.
http://www.youtube.com/watch?v=B8XaJbwwT68
Bob Clark
Robert Clark
At the beginning they are also moving around the engine which also
looks to be composite. The *overwhelmingly* key question about the
possibility of a reusable SSTO is whether or not it could be made
light enough. I'm arguing it can be if made of all composite
construction. This video gives further support of that argument.
Bob Clark
Pat Flannery
> At the beginning they are also moving around the engine which also
> looks to be composite. The *overwhelmingly* key question about the
> possibility of a reusable SSTO is whether or not it could be made
> light enough.
Although it could be made of some sort of RCC with a composite outer
shell, it might also be have a thin heat resistant metal internal liner
with carbon composite wrapping around it to take the pressure of the
combustion within it.
If it was all composite, its inner surface could probably get damaged
from the heat of the combustion process, meaning it would only be usable
for one flight.
Note that it it is heavy enough that they have to push it around on a
cart rather than carrying it, and that it takes some effort to move it
around, even on the cart.
From the amount of pushing they are doing, it looks like it weighs
between 500-1,000 pounds.
The composite crew/passenger compartment halves do look very light for
their size though compared to metal ones, as does the rest of the structure.
It would be interesting to know how many man-hours went into making the
parts and assembling them.
Pat
Greg D. Moore (Strider)
air-launched (which basically makes it 2 stages) suborbital craft further
that argument at all.
"Robert Clark" <rgrego...@yahoo.com> wrote in message
news:04d09b77-e0ad-47cb...@p35g2000yqh.googlegroups.com...
Robert Clark
The Air Force is researching reusable hydrocarbon-fueled first stage
boosters to be used with expendable upper stages to cut the costs to
space by 50%:
USAF Seeks Reusable Booster Ideas.
May 14, 2009
By Graham Warwick
"AFRL's reference concept includes an integral all-composite airframe
and tank structure that carries both internal pressure and external
flight loads. The concept vehicle is powered by pump-fed liquid-oxygen/
hydrocarbon rocket engines."
http://www.aviationweek.com/aw/generic/story_channel.jsp?channel=space&id=news/Reuse051409.xml
This article discusses wind tunnel tests of a scale-model of such a
booster:
AEDC team conducts first test on a reusable space plane.
Posted 12/16/2009 Updated 12/16/2009
http://www.arnold.af.mil/news/story.asp?id=123182588
It is interesting they are proposing an all-composite construction
including propellant tanks for this reusable hydrocarbon-fueled first
stage booster.
As I have argued, an all-composite, hydrocarbon-fueled design would
allow even a reusable single-stage-to-orbit vehicle.
Bob Clark
BradGuth
cyclopropane?
~ BG
Robert Clark
I had a debate with you a few months ago about which should be the
propellant to use for a single stage to orbit vehicle. I argued it
should be hydrogen/LOX because that gave the highest Isp and therefore
the lowest propellant mass. I also argued that both the DC-X
demonstrator and the proposed VentureStar vehicles used hydrogen fuel.
However, I now realize Isp is not the only key variable. There are
other variables that can overwhelm the Isp advantage of hydrogen/LOX.
Then propellant combinations with dense fuels and/or oxidizers,
including H2O2, may indeed be better suited for a SSTO vehicle.
Here's one report that discusses using H2O2 as the oxidizer for a
SSTO:
A Single Stage to Orbit Rocket with Non-Cryogenic Propellants.
Abstract
"Different propellant combinations for single-stage-to-orbit-rocket
applications were compared to oxygen/hydrogen, including nitrogen
tetroxide/hydrazine, oxygen/methane, oxygen/propane, oxygen/RP-1,
solid core nuclear/hydrogen, and hydrogen peroxide/JP-5. Results show
that hydrogen peroxide and JP-5, which have a specific impulse of 328
s in vacuum and a density of 1,330 kg/cu m. This high-density jet fuel
offers 1.79 times the payload specific energy of oxygen and hydrogen.
By catalytically decomposing the hydrogen peroxide to steam and oxygen
before injection into the thrust chamber, the JP-5 can be injected as
a liquid into a high-temperature gas flow. This would yield superior
combustion stability and permit easy throttling of the engine by
adjusting the amount of JP-5 in the mixture. It is concluded that
development of modern hydrogen peroxide/JP-5 engines, combined with
modern structural technology, could lead to a simple, robust, and
versatile single-stage-to-orbit capability."
http://www.erps.org/docs/SSTORwNCP.pdf [full text]
Bob Clark
Robert Clark
Spacefaring Civilization_ brought to mind a key advantage of this
reconfigured X-33/VentureStar that I hadn't considered before:
"The shuttle is a fiscal disaster not because it is reusable, but
because both its technical and programmatic bases are incorrect. The
shuttle is a partially reusable launch vehicle: Its lower stages are
expendable or semi-salvageable while the upper stage (the orbiter ) is
reusable. As aesthetically pleasing as this configuration may appear
to some, from an engineering point of view this is precisely the
opposite of the correct way to design a partially reusable launch
system. Instead, the lower stages should be reusable and the upper
stage expendable. Why? Becasue the lower stages of a multi-staged
booster are far more massive than the upper stage: so if only one or
the other is to be reusable, you save much more money by reusing the
lower stage. Furthermore, it is much easier to make the lower stage
reusable, since it does not fly as high or as fast, and thus takes
much less of a beating during reentry. Finally the negative payload
impact of adding those systems required for reusability is much less
if they are put on the lower stage than the upper. In a typical two-
stage to orbit system for example every kilogram of extra dry mass
added to the lower stage reduces the payload delivered to orbit by
about 0.1 kilograms, whereas a kilogram of extra dry mass on the upper
stage causes a full kilogram of payload loss. The Shuttle is actually
a 100-tonne to orbit booster, but because the upper stage is a
reusable orbiter vehicle with a dry mass of 80 tonnes, only 20 tonnes
of payload is actually delivered to orbit. From the amount of smoke,
fire, and thrust the Shuttle produces on the launch pad, it should
deliver five times the payload to orbit of a Titan IV, but because it
must launch the orbiter to space as well as the payload, its net
delivery capability only equals that of the Titan. There is no need
for 60-odd tonnes of wings, landing gear and thermal protection
systems in Earth orbit, but the shuttle drags them up there (at a cost
of $10 million per tonne) anyway each time it flies. In short the
Space Shuttle is so inefficient because *it is built upside down*."
{emphasis in the original.}
_Entering Space_, p. 29.
Zubrin makes a key point about that dry weight of 80,000 kg of the
orbiter, which is essentially an upper stage, that needs to be carried
along to bring that approx. 20,000 kg of payload to orbit. That 4 to 1
ratio of the upper stage dry weight to payload weight struck me
because the upper stage for rockets is usually a quite lightweight
structure. Then the shuttle is quite poor on this measurement. I then
thought of the reconfigured kerosene version of the VentureStar I was
considering and realized that it was actually quite good on this
scale. It could carry ca. 125,000 payload to orbit with a vehicle dry
weight ca. 82,000 kg.
Actually the total shuttle system as a whole is even worse on this
scale. This site gives the specifications for some launch systems:
Space Launch Report Library.
http://www.spacelaunchreport.com/library.html
Here's the page for the shuttle system:
Space Launch Report: Space Shuttle.
http://www.spacelaunchreport.com/sts.html
You can calculate the total dry weight by subtracting off the
propellant weight from the gross weight for each component. I
calculate a total dry weight of 310,850 kg to a payload weight of
24,400 kg, a ratio of 12.7 to 1. In contrast the reconfigured
VentureStar has this ratio going in the other direction, that is, the
payload weight is larger than the vehicle dry weight.
This is important because the total dry weight is a key parameter for
the cost of a launch system. I looked at some of the vehicles listed
on the Spacelaunchreport.com page and all the ones I looked had the
total dry weight higher than the payload weight. For instance for the
Delta IV, it's a dry weight of 37,780 kg to a payload weight of 8,450
kg, for a ratio of 4.5 to 1.
For the Atlas V 401 it was 25,660 kg dry weight to a 12,500 kg payload
weight, for a ratio of 2 to 1. This was actually one of the better
ones. All the ones I looked at, all had a total dry weight
significantly higher than the payload weight, usually at least by a 3
to 4 to 1 ratio.
Then the reconfigured VentureStar would be important in that it could
reverse this trend (perhaps for the first time?) in making the dry
weight actually less than the payload that could be lofted to orbit.
Note that not even the original, planned VentureStar could accomplish
this, having a dry weight of about 100,000 kg to a payload capacity of
20,000, a ratio of 5 to 1.
The reconfigured kerosene-fueled VentureStar would have a greater
propellant mass using dense propellants, but the propellant costs are
a relatively small proportion of the launch costs. The more important
parameter of dry weight would actually be less.
Note also that the reconfigured kerosene VentureStar could accomplish
this feat of having a higher payload capacity than its dry weight,
while having a payload capacity that would be close to or exceed that
of all the former or planned U.S. launchers, and while being of
significantly smaller dimensions. See the attached image drawn to
scale showing some key U.S. launchers compared to the VentureStar.
Note that despite its small size it would be carrying more payload
than the shuttle, the Ares I, the Saturn V and nearly that of the Ares
V.
Another factor that I somehow missed when I first wrote on this was
the great reduction in launch costs. I somehow didn't make the
connection between the increase in payload capacity over the original
VentureStar configuration and that of the kerosene fueled one.
The development costs for the VentureStar or any launch vehicle are
figured into the launch costs. Then the estimated per kg launch costs
of ca. $1,000/kg for the original VentureStar are based on the late
90's estimated development costs for the VentureStar. However, a big
part of that development cost was due to the composite design which
was significantly more expensive then than now. Recall the point I
made before about the reduction in costs of composite materials
leading to auto makers including them more and more in passenger cars,
and this reduction in cost makes them now economically feasible for an
all composite SSTO.
Also, hydrogen engines and associated systems are generally more
expensive than kerosene ones. So the reconfigured VentureStar would
have a lower cost on that component as well. Then the total
development cost even including inflation for the reconfigured
VentureStar might be at or even below that of the 1990's estimates for
the original hydrogen-fueled VentureStar. This means the per launch
costs of the new version should be at or below that of the original
version.
But the reconfigured VentureStar can carry 6 times the payload of the
original VentureStar! This means the per kg launch costs would be
1/6th as much or only $166 per kg! This is such an *extreme* reduction
in launch costs over the current costs, that the calculation I made
for how much you could reduce the weight of the propellant tanks has
to be done in a more serious fashion.
Note that all the other systems for the VentureStar were progressing
well. It was only the relatively trivial problem of not using a strong
enough glue for the composite propellant tanks, that led to the
program being canceled. Then this is so trivial compared to the
complexity of the other systems and the importance of having a fully
reusable launch system is so clear, that a better course would have
been to open up a competition to find ways of getting the composite
tanks to work.
I gave a few different possibilities for lightweighting the propellant
tanks in section II of the first post in this thread. A few were
theoretical, not being tried before. However, the one involving
partitioned tanks is just basic engineering so I'll present a detailed
calculation for that in the next post.
Bob Clark
VentureStar, Shuttle, Ares I-V, and Saturn V size comparison.
http://i49.tinypic.com/2z3rup1.jpg
Robert Clark
>Table of Contents.
>...
>II.)Lightweight propellant tanks.
>...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.
>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.
We'll view the X-33 hydrogen tanks standing vertically as conical
with flattened front and back. This report on page 19 by the PDF file
page numbering gives the dimensions of the X-33 hydrogen tanks as 28.5
feet long, 20 feet wide and 14 feet high:
Final Report of the X-33 Liquid Hydrogen Tank Test Investigation Team.
http://alpha.tamu.edu/public/jae/misc/tankreport.pdf
Call it 9 meters long, 6 meters wide, and 4.3 meters deep for this
calculation. I'll simplify the calculation by approximating the shape
as rectangular, i.e., uniformly 6 meters wide. See the attached image.
Note that the rounded portions of the sides, top and bottom will be
considered separately. I'll call the vertical length of each section
x, and the bulkhead thickness h. Since the length of the tank is 9m,
the number of sections is 9/x.
I'm doing the calculation for kerosene/LOX propellant tanks, but
approx. same size as the X-33 tanks. Typically these are pressurized
in the 20-40 psi range. I'll take it as 30 psi; call it 2 bar, 2x10^5
Pa. Referring to the drawing of the tank, each bulkhead takes part in
supporting the internal pressure of the two sections on either side of
it. This means for each section the internal pressure is supported by
one-half of each bulkhead on either side of it, which is equivalent to
saying each bulkhead supports the internal pressure of one section.
The force on each section is the cross-sectional area times the
internal pressure, so 6m*x*(2*10^5 Pa), with x as in the diagram the
vertical length of each section. The bulkhead cross-sectional area is
6m*h, with h the thickness of the bulkheads. Then the pressure the
bulkheads have to withstand is 6m*x*(2*10^5 Pa)/6m*h = (2*10^5 Pa)*x/
h.
The volume of each bulkhead is 6m*h*4.3m. The density of aluminum-
lithium alloy is somewhat less than aluminum, call it 2,600 kg/m^3. So
the mass of each bulkhead is (2,600 kg/m^3)*6m*h*4.3m = 67,080*h. Then
the total mass of all the 9/x bulkheads is (9/x)*67080*h = 603,720*(h/
x).
Note that additionally to the horizontal bulkheads shown there will be
vertical bulkheads on the sides. These will have less than 1/10 the
mass of the horizontal bulkheads because the length of each section x
will be small compared to the width of 6m, and will have likewise
small contribution to the support of the internal pressure.
The tensile strength of some high strength aluminum-lithium alloys can
reach 700 MPa, 7*10^8 Pa. Then the pressure the bulkheads are
subjected to has to be less than or equal to this: (2*10^5 Pa)*x/h <=
7*10^8 Pa, so x/h <= 3,500, and h/x => 1/3,500. Therefore the total
mass of the bulkheads = 603,720*(h/x) => 172.5 kg. Note we have not
said yet how thick the bulkheads have to be only that their total mass
is at or above 172.5 kg, for one of the twin rear tanks. It's twin
would also require 172.5 kg in bulkhead mass. The third, forward, tank
had about 2/3rds the volume of these twin rear tanks so I'll estimate
the bulkhead mass it will require as 2/3rds of 172.5 kg, 115 kg. Then
the total bulkhead mass would be 460 kg, about 15% of the 3,070 kg
tank mass I calculated for the reconfigured X-33.
This is for the bulkheads resisting the outwards pressure of the
sections. Notice I did not calculate the pressure inside the tank on
the bulkheads from the propellant on either side. This is because the
pressure will be equalized on either side of the bulkheads. However,
we will have to be concerned about the pressure on the rounded right
and left sides of the tank, and the rounded top and bottom of the
tank, where the pressure is not equalized on the outside of the tank.
Before we get to that, remember the purpose of partitioning the tank
was to minimize the bowing out of the front and back sides from the
internal pressure. Consider this page then that calculates the
deflection of a flat plat under a uniform load:
eFunda: Plate Calculator -- Clamped rectangular plate with uniformly
distributed loading.
[img]http://www.efunda.com/formulae/solid_mechanics/plates/images/
CCCC_PUniform.gif[/img]
"This calculator computes the maximum displacement and stress of a
clamped (fixed) rectangular plate under a uniformly distributed load."
http://www.efunda.com/formulae/solid_mechanics/plates/calculators/CCCC_PUniform.cfm
In the data input boxes, we'll put 200 kPa for the uniform load, 6
meters for the horizontal distance, .3 m, say, for the vertical
distance x, and 6 mm for the thickness of the plate. For the vertical
distance x I'm taking a value proportionally small compared to the
tank width, but which won't result in an inordinate number of
partitioned sections of the tank. For the thickness I'm taking a value
at 1/1000th the width of the tank, which is common for cylindrical
tanks. For the material specifications for aluminum-lithium we can
take the Young's modulus as 90 GPa. Then the calculator gives the
deflection as only 2.35mm, probably adequate.
However, we still have to consider what happens to the rounded sides
and the bottom and top. Look at the last figure on this page:
Thin-Walled Pressure Vessels.
[img]http://www.efunda.com/formulae/solid_mechanics/mat_mechanics/
images/PressureVesselCylindricalC.gif[/img]
http://www.efunda.com/formulae/solid_mechanics/mat_mechanics/pressure_vessel.cfm
It shows the calculation for the hoop stress of a cylindrical pressure
vessel. The calculation given is 2*s*t*dx = p*2*r*dx, using s for the
hoop stress. This implies, s = p*r/t, or equivalently t = p*r/s. So
for a given material strength s, the thickness will depend only on the
radius and internal pressure.
However, what's key here is the same argument will apply in the figure
if one of the sides shown is flat, instead of curved. Therefore in our
scenario, the rounded sides, top and bottom, which we regard as half-
cylinders, will only need the thickness corresponding to a cylinder of
their same diameter, i.e., one of a diameter of 4.3m.
So the rounded portions actually require a smaller thickness than what
would be needed for a cylinder of diameter of the full 6m width of the
tank.
This means the partitioned tank requires material of somewhat less
mass than a cylindrical tank of dimension the full width of the tank
plus about 15% of that mass as bulkheads.
Bob Clark
tank drawing.
[img]http://i49.tinypic.com/a9qy4w.jpg[/img]
Bob Myers
> expendable or semi-salvageable while the upper stage (the orbiter ) is
> reusable. As aesthetically pleasing as this configuration may appear
> to some, from an engineering point of view this is precisely the
> opposite of the correct way to design a partially reusable launch
> system. Instead, the lower stages should be reusable and the upper
> stage expendable. Why? Becasue the lower stages of a multi-staged
> booster are far more massive than the upper stage: so if only one or
> the other is to be reusable, you save much more money by reusing the
> lower stage.
I don't say whether Zubrin's conclusion is correct or not, but the
logic in the above works only if "far more massive" always
translates to "far more expensive." I don't believe that's necessarily
the case.
Bob M.
Pat Flannery
> Robert Clark wrote:
>> expendable or semi-salvageable while the upper stage (the orbiter ) is
>> reusable. As aesthetically pleasing as this configuration may appear
>> to some, from an engineering point of view this is precisely the
>> opposite of the correct way to design a partially reusable launch
>> system. Instead, the lower stages should be reusable and the upper
>> stage expendable. Why? Becasue the lower stages of a multi-staged
>> booster are far more massive than the upper stage: so if only one or
>> the other is to be reusable, you save much more money by reusing the
>> lower stage.
>
> I don't say whether Zubrin's conclusion is correct or not, but the
> logic in the above works only if "far more massive" always
> translates to "far more expensive." I don't believe that's necessarily
> the case.
>
>
There's a three-stage fully reusable Lockheed concept from the early
1960's on the bottom of this webpage, as well as a really bizarre
Aerojet-General flying wing reusable spacecraft from the same period
further up: http://dreamsofspace.nfshost.com/1965orbitingstations.htm
Pat
Robert Clark
specifications of some launch vehicles. Here's the page for the Falcon
1:
Space Launch Report: SpaceX Falcon Data Sheet.
http://www.spacelaunchreport.com/falcon.html
Quite interesting is that the total mass and dry mass values for the
Falcon 1 first stage with Merlin 1C engine give a mass ratio of about
20 to 1. This is notable because a 20 to 1 mass ratio is the value
usually given for a kerosene-fueled vehicle to be SSTO. However, this
is for the engine having high vacuum Isp ca. 350 s. The Merlin 1C with
a vacuum Isp of 304 s probably wouldn't work.
However, there are some high performance Russian kerosene engines that
could work. Some possibilities:
Engine Model: RD-120M.
http://www.astronautix.com/engines/rd120.htm#RD-120M
RD-0124.
http://www.astronautix.com/engines/rd0124.htm
Engine Model: RD-0234-HC.
http://www.astronautix.com/engines/rd0234.htm
However, I don't know if this third one was actually built, being a
modification of another engine that burned aerozine.
Some other possibilities can be found on the Astronautix site:
Lox/Kerosene.
http://www.astronautix.com/props/loxosene.htm
And on this list of Russian rocket engines:
Russian/Ukrainian space-rocket and missile liquid-propellant engines.
http://www.b14643.de/Spacerockets_1/Diverse/Russian%20engines/engines.htm
The problem is the engine has to have good Isp as well as a good T/W
ratio for this SSTO application. There are some engines listed that
even have a vacuum Isp above 360 s. However, these generally are the
small engines used for example as reaction control thrusters in orbit
and usually have poor T/W ratios.
For the required delta-V I'll use the fact that a dense propellant
vehicle may only require a delta-V of 8,900 m/s, compared to a
hydrogen-fueled vehicle which may require in the range of 9,100 to
9,200 m/s. The reason for this is explained here:
Hydrogen delta-V.
http://yarchive.net/space/rocket/fuels/hydrogen_deltav.html
Then when you add on the fact that launching near the equator gives
you 462 m/s for free from the Earth's rotation, we can take the
required delta-V that has to be supplied by the kerosene-fueled
vehicle as 8,500 m/s.
I'll focus on the RD-0124 because of its high Isp, 359 s vacuum and
331 s sea level. On the "Russian/Ukrainian space-rocket and missile
liquid-propellant engines" page its sea level thrust is given as
253,200 N, 25,840 kgf. However, the Falcon 1 first stage weighs 28,553
kg. So we'll need two of them. Each weighs 480 kg, so two would be 960
kg. This is 300 kg more than the single Merlin 1C. So the dry mass of
the Falcon 1 first stage is raised to 1,751 kg. There is a RD-0124M
listed on the Astronautix page that only weighs 360 kg, but its sea
level Isp and thrust are not given, so we'll use the RD-0124 until
further info on the RD-0124M is available.
Taking the midpoint value of the Isp as 345 s we get a delta-V of
345*9.8ln(1 + 27102/1751) = 9,474 m/s (!) Note also the achieved delta-
V would actually be higher than this because the trajectory averaged
Isp is closer to the vacuum value since the rocket spends most of the
time at altitude.
This calculation did not include the nose cone fairing weight of 136
kg. However, the dry mass for the first stage probably includes the
interstage weight, which is not listed, since this remains behind with
the first stage when the second stage fires. Note then that the
interstage would be removed for the SSTO application. From looking at
the images of the Falcon 1, the size of the cylindrical interstage in
comparison to the conical nose cone fairing suggests the interstage
should weigh more. So I'll keep the dry mass as 1,751 kg.
Now considering that we only need 8,500 m/s delta-V we can add 636 kg
of payload. But this is even higher than the payload capacity of the
two stage Falcon 1!
We saw that the thrust value of the RD-0124 is not much smaller than
the gross weight of the Falcon 1 first stage. So we can get a vehicle
capable of being lifted by a single RD-0124 by reducing the propellant
somewhat, say by 25%. This reduces the dry weight now since one
RD-0124 weighs less than a Merlin 1C and the tank mass would also be
reduced 25%. Using an analogous calculation as before, the payload
capacity of this SSTO would be in the range of 500 kg.
We can perform a similar analysis on the Falcon 1e first stage that
uses the upgraded Merlin 1C+ engine. Assuming the T/W ratio of the
Merlin 1C+ is the same as that of the Merlin 1C, the mass of the two
of the RD-124's would now be only 100 kg more than the Merlin 1C+.
The dry mass and total mass numbers on the SpaceLaunchReport page for
the Falcon 1e are estimated. But accepting these values we would be
able to get a payload in the range of 1,800 kg. This is again higher
than the payload capacity of the original two stage Falcon 1e. In fact
it could place into orbit the 1-man Mercury capsule.
The launch cost of the Falcon 1, Falcon 1e is only about $8 million -
$9 million. So we could have the first stage for that amount or
perhaps less since we don't need the engines which make up the bulk of
the cost. How much could we buy the Russian engines for? This article
says the much higher thrust RD-180 cost $10 million:
From Russia, With 1 Million Pounds of Thrust.
Why the workhorse RD-180 may be the future of US rocketry.
Issue 9.12 | Dec 2001
"This engine cost $10 million and produces almost 1 million pounds of
thrust. You can't do that with an American-made engine."
http://www.wired.com/wired/archive/9.12/rd-180.html
This report gives the price of the also much higher thrust AJ26-60,
derived from the Russian NK-43, as $4 milliion:
A Study of Air Launch Methods for RLVs.
Marti Sarigul-Klijn, Ph.D. and Nesrin Sarigul-Klijn, Ph.D.
AIAA 2001-4619
"The main engine is currently proposed as the 3,260
lb. RP-LOX Aerojet AJ26-60, which is the former
Russian NK-43 engine. Thrust to weight of 122 to
1 compares to the Space Shuttle Main Engine’s
(SSME) 67 to 1 and specific impulse (Isp = 348.3
seconds vacuum) is 50 to 60 seconds better than
the Atlas II, Delta II, or Delta III RP-LOX engines.
A total of 831 engines have been tested for
194,000 seconds. These engines are available for
$4 million each, which is about 10% the cost of a
SSME."
http://mae.ucdavis.edu/faculty/sarigul/aiaa2001-4619.pdf
Then the much lower thrust RD-0124 could quite likely be purchased
for less than $4 million. So the single RD-0124 powered SSTO could be
purchased for less than $12 million.
Even though the mathematics says it should be possible, and has been
for decades, it is still commonly believed that SSTO performance with
chemical propulsion is not possible even among experts in the space
industry:
Space Tourism is a Hoax
By Fredrick Engstrom and Heinz Pfeffer
11/16/09 09:02 AM ET
"In 1903, the Russian scientist Konstantin Tsiolkovsky established the
so-called rocket equation, which calculates the initial mass of a
rocket needed to put a certain payload into orbit, given that the
orbital speed is fixed at 28,000 kilometers per hour, and that the
maximum speed of the gas exhausted from the rocket that propels it
forward is also fixed.
"You quickly find that the structure and the tanks needed to contain
the fuel are so heavy that you will never be able to orbit a
significant payload with a single-stage rocket. Thus, it is necessary
to use several rocket stages that are dumped on the way up to get any
net mass, i.e. payload, into orbit.
"Let us look at the most successful rocket on the market — the
European Ariane 5. Its start weight is 750 tons, of which 650 tons are
fuel, 80 tons are structure and around 20 tons are left for low Earth
orbit payload.
"You can have a different number of stages, and you can look for minor
improvements, but you can never get around the fact that you need big
machines that are staged to reach orbital speed. Not much has happened
in propulsion in a fundamental sense since Wernher von Braun’s Saturn
rocket. And there is nothing on the horizon, if you discount
controlling gravity or some exotic technology like that. In any case,
it is not for tomorrow."
http://www.spacenews.com/commentaries/091116-space-tourism-hoax.html
The Cold Equations Of Spaceflight.
by Jeffrey F. Bell
Honolulu HI (SPX) Sep 09, 2005
"Why isn't Mike Griffin pulling out the blueprints for X-30/NASP, DC-X/
Delta Clipper, or X-33/VentureStar? Billions of dollars were spent on
these programs before they were cancelled. Why aren't we using all
that research to design a cheap, reusable, Single-Stage-To-Orbit
vehicle that operates just like an airplane and doesn't fall in the
ocean after one flight?"
"The answer to this question is: All of these vehicles were fantasy
projects. They violated basic laws of physics and engineering. They
were impossible with current technology, or any technology we can
afford to develop on the timescale and budgets available to NASA. They
were doomed attempts to avoid the Cold Equations of Spaceflight."
http://www.spacedaily.com/news/oped-05zy.html
Then it is important that such a SSTO vehicle be produced even if
first expendable to remove the psychological barrier that it can not
be done. Once it is seen that it can be done, and in fact how easily
and cheaply it can be done, then there it will be seen that in fact
the production of SSTO vehicles are really no more difficult than
those of multistage vehicles.
Then will be opened the floodgates to reusable SSTO vehicles, and low
cost passenger space access as commonplace as trans-oceanic air
travel.
Bob Clark
Me
> Then it is important that such a SSTO vehicle be produced even if
> first expendable to remove the psychological barrier that it can not
> be done. Once it is seen that it can be done, and in fact how easily
> and cheaply it can be done, then there it will be seen that in fact
> the production of SSTO vehicles are really no more difficult than
> those of multistage vehicles.
> Then will be opened the floodgates to reusable SSTO vehicles, and low
> cost passenger space access as commonplace as trans-oceanic air
> travel.
>
More clueless BS. Clark thinks he is smarter than everyone else.
This is a sign of a mental problem.
Robert Clark
> For the required delta-V I'll use the fact that a dense propellant
> vehicle may only require a delta-V of 8,900 m/s, compared to a
> hydrogen-fueled vehicle which may require in the range of 9,100 to
> 9,200 m/s. The reason for this is explained here:
>
> Hydrogen delta-V.http://yarchive.net/space/rocket/fuels/hydrogen_deltav.html
This is another key advantage of dense propellant vehicles for the
SSTO application, that the delta-V to orbit would be about 300 m/s
less than for a hydrogen-fueled SSTO vehicle. The main idea behind
this is that dense propellant vehicles burn mass so much more quickly
that they achieve the speed needed to attain the right altitude for
orbit more quickly. Since the gravity loss is dependent on the time
spent on this vertical portion of the trip, dense propellant vehicles
experience less gravity loss. Still, the explanation is probably not
easy to grasp unless you do the actual numerical calculations over the
trajectory of the flight. However, I can show an approximate
calculation that makes the idea more understandable below.
This Wikipedia article also mentions the fact that dense propellant
vehicles require 300 m/s less delta-V to orbit than hydrogen vehicles:
Single-stage-to-orbit.
# 4 Dense versus hydrogen fuels.
"The end result is the thrust/weight ratio of hydrogen-fueled engines
is
30–50% lower than comparable engines using denser fuels."
"This inefficiency indirectly affects gravity losses as well; the
vehicle has
to hold itself up on rocket power until it reaches orbit. The lower
excess
thrust of the hydrogen engines due to the lower thrust/weight ratio
means
that the vehicle must ascend more steeply, and so less thrust acts
horizontally.
Less horizontal thrust results in taking longer to reach orbit, and
gravity
losses are increased by at least 300 meters per second. While not
appearing
large, the mass ratio to delta-v curve is very steep to reach orbit in
a
single stage, and this makes a 10% difference to the mass ratio on top
of the
tankage and pump savings."
http://en.wikipedia.org/wiki/Single-stage-to-orbit#Dense_versus_hydrogen_fuels
However, the explanation given here is not quite correct. This rather
implies
it is a function of greater thrust/weight ratio only. But in actual
fact
the lowered delta-V required for dense fuels applies *even when the
hydrogen
and the dense fuel vehicles have the same thrust/weight ratio*.
For the calculation of the delta-V savings for dense fuels, suppose
both the
dense-fueled and hydrogen-fueled vehicles have a initial T/W of, say,
1.3. Let Mi
be the initial gross mass of the vehicle, r the constant propellant
flow rate, Ve the
exhaust velocity, a(t) the acceleration, changing with time, of the
vehicle due to
the thrust, and g the acceleration due to gravity 9.8 m/s^2. Then the
mass of the
vehicle at time t is Mi-r*t, and the thrust force is (Mi-r*t)a(t).
We'll use the fact that the thrust of a rocket is (propellant flow
rate)x(exhaust velocity) to get the equation (Mi-r*t)a(t) = r*Ve. We
can solve this for the acceleration to get a(t) = r*Ve/(Mi-r*t).
Now because we set the initial thrust/weight ratio as 1.3 we know
that thrust = r*Ve = 1.3(g*Mi), so Mi = r*Ve/(1.3g). Then plug this
into the equation for acceleration to get: a(t) = r*Ve/(r*Ve/(1.3g) -
r*t) = Ve/(Ve/(1.3.g) - t). Quite notable here is that the propellant
flow rate cancels out and the acceleration due to the thrust depends
only on the exhaust velocity Ve, or equivalently, only on the Isp.
Then for the vertical portion of the trip where gravity drag takes
place, the rocket's
acceleration will be Ve/(Ve/(1.3g)-t) - 9.8. Now it may not be
apparent at first glance but this formula says the acceleration is
greater for a smaller exhaust velocity Ve, so for a smaller Isp. To
make it clearer multiply top and bottom of the expression for a(t) by
1.3g to bring it to 1.3g*Ve/(Ve-1.3g*t). Then if you do the division
this becomes 1.3g + t*(1.3g)^2/(Ve-1.3g*t). Now you see because the Ve
is in the denominator the expression is larger when Ve is *smaller*.
So a dense propellant with a lower Isp will accelerate faster during
this vertical portion of the trip meaning it spends less time when
gravity drag is operating so that gravity drag is reduced. You
couldn't make the Isp be arbitrarily small though because that would
result in huge fuel loads and tanks, and, most importantly, engines to
get the vehicle off the ground.
Bob Clark
Robert Clark
kerosene would allow the hydrogen-fueled suborbital X-33 to now become
an orbital craft. However, I thought it would be able to carry minimal
payload if any.
However, I realize I used too low a value for the density of chilled
LOX at 1,160 kg/m^3. It should be actually about 10% higher than the
usual 1,142 kg/m^3.
This is described here:
Alternate Propellants for SSTO Launchers.
Dr. Bruce Dunn
Adapted from a Presentation at:
Space Access 96
Phoenix Arizona
April 25 - 27, 1996
http://www.dunnspace.com/alternate_ssto_propellants.htm
In table 2 it gives the densities of some chilled fuels including
kerosene, i.e., RP-1, and of LOX. The density given for the chilled
kerosene is 867 kg/m^3, and for LOX 1,262 kg/m^3. So for the 296 m^3
volume I was taking for the X-33 propellant tanks and a 2.7 mixture
ratio for the NK-33 engine, this gives a kero/LOX propellant mass of
332,600 kg.
Now taking the average Isp of the NK-33 as 315 s, this gives a delta-V
for the 21,700 kg dry mass, reconfigured X-33 of 8,797 m/s. But when
you take into account you get a 462 m/s velocity boost for free from
launching at the equator, you only need about 8,500 m/s delta-V to be
provided by the rocket to reach orbit.
This allows us to add payload. Adding 2,300 kg payload, the delta-V
becomes 8,500 m/s, sufficient for orbit. We can actually get higher
payload than this by using more energetic hydrocarbons than kerosene.
For instance in table 2 of Dunn's paper on alternate SSTO propellants,
he gives the payload for chilled methylacetylene/LOX as 24% higher
than for chilled kero/LOX. This would be a payload of 2,850 kg.
These payload amounts would also allow the X-33 to carry a 2 man crew
in its 5 by 10 foot payload bay in a tandem arrangement a la the F-14
seating arrangement.
So you could get a fully reusable, SSTO vehicle at much reduced price
than the full-sized VentureStar. This article gives the price to build
a new X-33 as $360 million in 1998 dollars:
Adventure star.
http://www.flightglobal.com/pdfarchive/view/1998/1998%20-%203141.html
Even taking into account inflation, the cost to build the kerosene-
fueled version should be comparable or perhaps even less because of
the drop in prices for carbon composites and because kerosene engines
are generally cheaper than hydrogen ones.
The launch preparation costs should also be low since the X-33 was
expected to be operated by only a 50 man ground crew compared to the
18,000 required for the shuttle system:
Lockheed Secret Projects: Inside the Skunk Works.
By Dennis R. Jenkins
http://books.google.com/books?id=DUkl5bH6k6EC&lpg=PA95&dq=x-33%20venturestar&lr=&pg=PA106#v=onepage&q=&f=true
Say the builder expected 25% profit over costs of the vehicle over 100
flights. That would be a charge of $4.5 million per flight. At a 2,850
kg payload capacity that would be $1,580 per kilo, or $720 per pound,
to orbit. Not as good as the full-sized VentureStar but still
significantly better than current launch prices.
Note that the other half-scale suborbital demonstrators for the NASA
RLV program by Rockwell and McDonnell-Douglas (see images linked
below) could be built for comparable prices and would likewise become
full orbital craft by switching to kerosene or other dense propellant.
Then we could have 3 separate designs for fully reusable SSTO vehicles
at costs that could allow fully private financing that would
significantly reduce launch costs and would allow manned flights.
Successful operation of these X-33-sized orbital vehicles at a profit
would encourage private financing to build the full-scale VentureStar-
sized RLV's that could bring launch costs down to the $100 to $200 per
kilo range.
Bob Clark
Robert Clark
[re-posted to correct typos.]
No. I'm reporting what some experts in the field have said, that it
is easier to produce a SSTO vehicle with dense fuels rather than with
hydrogen.
Some examples:
Single Stage To Orbit Mass Budgets Derived From Propellant Density and
Specific Impulse.
John C. Whitehead, Lawrence Livermore National Laboratory.
32nd AIAA/ASME/SAE/ASEE Joint Propulsion Conference.
Lake Buena Vista, FL July 1-3, 1996
Abstract
"The trade between specific impulse and density is examined
in view of SSTO requirements. Mass allocations for
vehicle hardware are derived from these two properties, far
several propellant combinations and a dual-fuel case. This
comparative analysis, based on flight-proven hardware,
indicates that the higher density of several alternative
propellants compensates for reduced Isp, when compared
with cryogenic oxygen and hydrogen. Approximately half
the orbiting mass of a rocket-propelled SSTO vehicle must
be allocated to propulsion hardware and residuals. Using
hydrogen as the only fuel requires a slightly greater fraction
of orbiting mass for propulsion, because hydrogen engines
and tanks are heavier than those for denser fuels. The
advantage of burning both a dense fuel and hydrogen in
succession depends strongly on tripropellant engine weight.
The implications of the calculations for SSTO vehicle
design are discussed, especially with regard to the necessity
to minimize non-tankage structure."
http://www.osti.gov/bridge/servlets/purl/379977-2LwFyZ/webviewable/379977.pdf
A Single Stage to Orbit Rocket with Non-Cryogenic Propellants.
Clapp, Mitchell B.; Hunter, Maxwell W.
AIAA, SAE, ASME, and ASEE, Joint Propulsion Conference and Exhibit,
29th, Monterey, CA, June 28-30, 1993.
Abstract
"Different propellant combinations for single-stage-to-orbit-rocket
applications were compared to oxygen/hydrogen, including nitrogen
tetroxide/hydrazine, oxygen/methane, oxygen/propane, oxygen/RP-1,
solid core nuclear/hydrogen, and hydrogen peroxide/JP-5. Results show
that hydrogen peroxide and JP-5, which have a specific impulse of 328
s in vacuum and a density of 1,330 kg/cu m. This high-density jet fuel
offers 1.79 times the payload specific energy of oxygen and hydrogen.
By catalytically decomposing the hydrogen peroxide to steam and oxygen
before injection into the thrust chamber, the JP-5 can be injected as
a liquid into a high-temperature gas flow. This would yield superior
combustion stability and permit easy throttling of the engine by
adjusting the amount of JP-5 in the mixture. It is concluded that
development of modern hydrogen peroxide/JP-5 engines, combined with
modern structural technology, could lead to a simple, robust, and
versatile single-stage-to-orbit capability."
http://www.erps.org/docs/SSTORwNCP.pdf
Alternate Propellants for SSTO Launchers.
Dr. Bruce Dunn
Adapted from a Presentation at:
Space Access 96
Phoenix, Arizona
April 25 =96 27, 1996
Introduction
"The most commonly proposed propellant combination for an SSTO
launcher is liquid oxygen and liquid hydrogen, at a mixture ratio of
approximately 6.0. There have been a number of studies of alternate
fuels for SSTO launchers, but they have been limited. To date, most
studies have concentrated on methane, propane and RP-1 burned with
liquid oxygen to the exclusion of other oxidizers and other fuels.
These studies have often, but not always shown lower vehicle dry
masses for hydrocarbon propellants (for the same payload size). The
lowest dry masses of all are found in dual-fuel vehicles, using dense
hydrocarbons early in the flight and hydrogen late in the ascent.
These vehicles however suffer from mechanical and structural
complexity over their single-fuel cousins, and are unlikely to
represent the least expensive way to get a defined payload to orbit."
http://www.dunnspace.com/alternate_ssto_propellants.htm
This is certainly a minority opinion that dense fuels are better for a
SSTO than hydrogen, but it has occurred numerous times in science that
the minority opinion turns out to be the correct one.
The argument for why dense propellants are better for a SSTO is quite
simple and can be understood by anyone familiar with the "rocket
equation" that describes the relationship between the exhaust
velocity and the mass of propellant for a rocket. Indeed the argument
is as about as close to a mathematical proof as you can get in
engineering.
First two key facts have to be kept in mind: 1.) the tank mass scales
by volume, *NOT* by the mass of the fluid contained. This means that
the same size and *same mass* tanks can hold about 3 times as much
kero/LOX as LH2/LOX. This is extremely important because the
propellant tanks make up the single biggest component of the dry
weight of a rocket, typically 30% to 40%, even more than that of the
engines.
And 2.) dense propellant engines such as kerosene ones typically have
thrust/weight ratios twice as good as hydrogen ones. This is key
because switching to kerosene means your fuel load and therefore gross
mass will be greater. But because of the kerosene engines better T/W
ratio, the increase in engine weight will be relatively small.
Many people get the second of these points. It's the reason why first
stages generally use kerosene or other dense propellant for example.
However, the first point most people are not as familiar with. But
it's the more important of the two because the increase in propellant
being carried far exceeds the increase needed to overcome the lowered
Isp of the dense propellants.
To see why tank mass scales with volume, take a look at the equations
for tank mass here:
Pressure vessel.
http://en.wikipedia.org/wiki/Pressure_vessel#Scaling
Note it depends only on tank dimensions, internal pressure, and
strength and density of the tank material. Then because the internal
pressure of the tanks will be about the same for the hydrogen case as
for the kerosene case, for proper operation of the turbopumps, the
kerosene filled tanks will hold about 3 times more propellant at the
same size and weight of the tanks.
Now for the calculation that switching to kerosene can result in
multiple times greater payload. The vacuum Isp for good hydrogen
engines is about 450 s, and for good kerosene ones about 350 s. This
means the mass ratio for a hydrogen SSTO is about 10 and for a
kerosene one it's about 20. These values are higher than what you
would expect based just on the vacuum Isp alone because you also have
to consider gravity and air drag, and the fact that the Isp is
decreased at sea level and low altitude.
Now suppose we switch our hydrogen-fueled SSTO for a kerosene-one
using the same sized tanks. The volume stays the same so the mass of
the tanks stays the same. But the amount of propellant is now about 3
times larger.
For the engines, since propellant mass makes up almost all the
vehicle gross weight, the gross weight will be about 3 times larger
too. So the engines will need about 3 times the thrust.
For the original hydrogen-engines the thrust/weight ratio was about
50 to 1. And since the gross mass was about 10 times the dry mass for
the hydrogen vehicle, this means the engine mass was about 1/5, or
20%, of the dry weight.
Now switching to kerosene makes the gross weight about 3 times
larger. If the kerosene engines had only a 50 to 1 T/W ratio then you
would need 3 times heavier engines so they would be at 3/5 of the dry
weight. But since the thrust/weight of the kerosene engines is twice
that of the hydrogen ones, the engine weight is 1.5/5, 30%, of the dry
weight so the vehicle dry weight is increased only by 10%, due to the
heavier engines.
Now since the mass ratio is 10 for the hydrogen case but 20 for the
kerosene, you normally need about twice the kerosene propellant for
the same sized vehicle+payload total to reach orbit. But what we
actually have is about 3 times more propellant in our kerosene
vehicle, 1.5 times more than is necessary to get the same vehicle size
and payload to orbit. The vehicle does weigh about 10% more in dry
weight, so then the total vehicle+payload weight that can now
be lifted to orbit will be 1.5/1.1 = 1.364 times higher than for the
hydrogen case.
Now for the hydrogen powered SSTO vehicles that have been proposed
the payload is a fraction of the vehicle dry weight. The 100,000 kg
dry weight of the VentureStar compared to the 20,000 kg payload
capacity is typical. Then the kerosene version of such a vehicle could
loft (1.364)*(120,000 kg) = 164,000 kg to orbit. Or considering that
our vehicle is at a dry weight of 110,000 kg with the kerosene-engine
change, the payload would be 54,000 kg, 2.7 times the payload weight
of
the hydrogen case.
As I said this is an easy calculation to do. But many people simply
won't do it. They have been so conditioned to think that Isp is the
most important thing that the assumption is hydrogen must be used for
an SSTO. It probably doesn't help matters the fact that the gross mass
becomes about 3 times as great with the dense propellants. Gross mass
has been frequently used as the measure of the cost of a launch
vehicle, which I like to call "the hegemony of the GLOW weight".
But this is actually a very poor measure to use. The reason is
propellant cost is a trivial component of the launch cost to orbit.
More important is the dry mass and complexity of the launch vehicle
for the payload that can be orbited. Then what's important is
switching to a dense propellant allows multiple times greater payload
at the same sized and similarly dry-massed vehicle.
Bob Clark
hal...@aol.com
airplane to get the in orbit portion to at least 50,000 feet above
most of the atmosphere, not tied to a single launch location, fly the
airplane to a convenient launch location, fuel to get to 50,000 feet
can be from tanker refueling along the way..........
granted for a really large payload a BIG HUGGER AIRLINER might need to
be a custom build, but the upsides are huge.
no risky loaded bomb launch being the first.
SSTO is just a distraction from the more important......
LOW COST TO ORBIT!!
J. Clarke
Why do people think that launching from 50,000 feet will help somehow?
Going into orbit is not a matter of going high, it's a matter of going
_fast_. Launching from 50,000 feet or from sea level you still need to
impart 18,000 miles an hour of delta-v. That's the hard part.
hal...@aol.com
the hardest part is having enough fuel onboard to get you thru the
dense lower atmosphere.those large tanks weigh more.
with a aircraft first stage that part is taken care of by a mature
well understood technology, and since in air refueling to release
altitude would be used lots of unnecessary mass wouldnt need lifited
off the pad..
plus the aircraft with space plane could be released at the equator
gaing some margins too. and bad weather would be much less of a issue.
no more storm clouds nearing pad troubles. just fly a few hundred
miles away to a nice clear area.
its far easier to accelerate a lower mass object, the
Greg D. Moore (Strider)
Because 50,000 feet gets you above the bulk of the atmosphere which provides
a decent bonus.
--
Greg Moore
Ask me about lily, an RPI based CMC.
Marvin the Martian
> On Mar 21, 11:49�am, "J. Clarke" <jclarke.use...@cox.net> wrote:
>> On 3/21/2010 10:42 AM, hall...@aol.com wrote:
>>
>> > Why take along EVERYTHING for a SSTO when the vehicle could use a
>> > airplane to get the in orbit portion to at least 50,000 feet above
>> > most of the atmosphere, not tied to a single launch location, fly the
>> > airplane to a convenient launch location, fuel to get to 50,000 feet
>> > can be from tanker refueling along the way..........
>>
>> > granted for a really large payload a BIG HUGGER AIRLINER might need
>> > to be a custom build, but the upsides are huge.
>>
>> > no risky loaded bomb launch being the first.
>>
>> > SSTO is just a distraction from the more important......
>>
>> > LOW COST TO ORBIT!!
>>
>> Why do people think that launching from 50,000 feet will help somehow?
>> Going into orbit is not a matter of going high, it's a matter of going
>> _fast_. �Launching from 50,000 feet or from sea level you still need to
>> impart 18,000 miles an hour of delta-v. �That's the hard part.
>
> the hardest part is having enough fuel onboard to get you thru the dense
> lower atmosphere.those large tanks weigh more.
It is apparent you're not acquainted with rocket science. Getting through
the "dense lower atmosphere" is no big deal. Von Braun did that with an
single stage alcohol fueled rocket 65 years ago.
The problem is getting up to orbital velocity.
Pat Flannery
>
> It is apparent you're not acquainted with rocket science. Getting through
> the "dense lower atmosphere" is no big deal. Von Braun did that with an
> single stage alcohol fueled rocket 65 years ago.
>
> The problem is getting up to orbital velocity.
If you can put the LOX aboard the rocket at altitude, where the humidity
is very low, you can eliminate the weight and complexity of having to
put insulation on the outside of the oxidizer tank section, as ice won't
form on it like it would if it were fueled and launched from the
surface. Not only does the booster then end up carrying the weight of
ice still sticking to it during ascent, but the ice that sheds can
damage the booster due to its mass and impact speed.
Pat
Jeff Findley
"Greg D. Moore (Strider)" <mooregr_d...@greenms.com> wrote in message
news:X6-dnTN9SfCfDzvW...@earthlink.com...
> J. Clarke wrote:
>> Why do people think that launching from 50,000 feet will help somehow?
>> Going into orbit is not a matter of going high, it's a matter of going
>> _fast_. Launching from 50,000 feet or from sea level you still need
>> to impart 18,000 miles an hour of delta-v. That's the hard part.
>
> Because 50,000 feet gets you above the bulk of the atmosphere which
> provides a decent bonus.
Specifically, you can optimize your engines for the much lower pressure of
50,000 feet (to vacuum), as opposed to the compromises necessary to make
them run at sea level.
Jeff
--
"Take heart amid the deepening gloom
that your dog is finally getting enough cheese" - Deteriorata - National
Lampoon
J. Clarke
> "Greg D. Moore (Strider)"<mooregr_d...@greenms.com> wrote in message
> news:X6-dnTN9SfCfDzvW...@earthlink.com...
>> J. Clarke wrote:
>>> Why do people think that launching from 50,000 feet will help somehow?
>>> Going into orbit is not a matter of going high, it's a matter of going
>>> _fast_. Launching from 50,000 feet or from sea level you still need
>>> to impart 18,000 miles an hour of delta-v. That's the hard part.
>>
>> Because 50,000 feet gets you above the bulk of the atmosphere which
>> provides a decent bonus.
>
> Specifically, you can optimize your engines for the much lower pressure of
> 50,000 feet (to vacuum), as opposed to the compromises necessary to make
> them run at sea level.
So how much do you think this gains you?
J. Clarke
So how much "weight and complexity" is involved with a little bit of
spray-on foam? And in practical terms how much difference is this going
to make? I'm sorry, but you're trying to reduce launch costs by
tackling an at best second order effect without dealing with the major
cost drivers. In any case the tankage on the X-33 is does not have
surfaces exposed to the airflow so this becomes a non-issue.
And if you're talking an X-33 it has to have a thermal protection system
for reentry anyway.
And the X-33 could not achieve more than half of orbital velocity on
HYDROGEN so how in the Hell do you expect it to do that with kerosene?
SSTO, if it can be done at all with chemical fuels, is _barely_ doable.
hal...@aol.com
>
> - Show quoted text -
If you call the airplane a non stage since it basically flies up to
release altitude then flies back to base.
A SSTO where the only stage is a orbital one is very doable.
espically since you dont have to carry ALL the fuel from the launch
pad to orbit.
with in flight refueling along the way it is a real winner.
no loaded bomb launch either:)
J. Clarke
What does that sentence mean? If it is single stage to orbit then there
is only one stage and since it achieves orbit it is necessarily "orbital".
But your assertion does not convince. You are posting on the Internet.
Most people posting on the Internet have opinions. Most of those
opinions are ignorant twaddle. So one must take your opinion as
ignorant twaddle until you can provide some numbers to go with it.
> espically since you dont have to carry ALL the fuel from the launch
> pad to orbit.
So where do you carry it? Is Spock beaming it into your vehicle with
the transporter or something?
> with in flight refueling along the way it is a real winner.
So how do you refuel it in flight?
> no loaded bomb launch either:)
So when does the "bomb" get "loaded" and how does that happen?
Show me the numbers on your airliner-launched SSTO. All that your
airliner brings to the party is a portable launch pad. Its effect on
the performance requirements is negligible.
Pat Flannery
> And the X-33 could not achieve more than half of orbital velocity on
> HYDROGEN so how in the Hell do you expect it to do that with kerosene?
X-33 was never designed to achieve orbital velocity, any more than DC-X
was; both were subscale proof-of-concept vehicles to try out engine,
aerodynamic, structure, and landing concepts.
If NASA and hadn't gotten so fixated on SSTO as a Shuttle replacement,
They could have built the Lockheed Starclipper concept, which would have
been a major step forward from the Shuttle as it eliminated the need for
the SRB's: http://www.astronautix.com/lvs/staipper.htm
They maybe even could have redesigned the VentureStar into a version
with drop tanks, as it owed a lot of its basic aerodynamics to
Starclipper, as well as the use of the (then classified) Starclipper's
linear plug-nozzle engine. Any performance shortfall generated by having
to switch to a aluminum-lithium LH2 tank from the composite one could
have been more than redressed by adding drop tanks to the design. The
advantages of high altitude fueling and launch for a Shuttle type
vehicle to avoid ice buildup on tankage using any sort of cryogenic
propellants go clean back to the Air Force/DARPA ALSV concept that
Dwayne Day is following the history of in The Space Review:
http://www.thespacereview.com/article/1569/1
http://www.thespacereview.com/article/1580/1
http://www.thespacereview.com/article/1591/1
Fluorine-deuterium? Oh, that will be cheap to use as fuel. ;-)
Pat
hal...@aol.com
>
> - Show quoted text -
Lets make it SIMPLE for you....
A large airliner with little fuel takes off, low fuel level keeps take
off weight down:)
with multiple in flight refuels, done every day in the military:) gets
the vehicle to near release altitude.
at this point the airliner sets off its afterburners and releases the
actual rocket stage, which achieves orbit.
the airliner flies back to base 100s if not a 1000 miles away.
a fully fuled rocket sitting on the pad is basically a loaded bomb.
a airliner launched rocket stage can use ejection seats for the
airliners crew, and a capsule safety pod for the rocket stage crew.
think out of the box, the box isnt your friend..............
Pat Flannery
> But your assertion does not convince. You are posting on the Internet.
> Most people posting on the Internet have opinions. Most of those
> opinions are ignorant twaddle. So one must take your opinion as ignorant
> twaddle until you can provide some numbers to go with it.
Yeah...but you are posting on the internet also, and I'm not seeing any
numbers so far. :-D
Pat
Jeff Findley
"J. Clarke" <jclarke...@cox.net> wrote in message
news:ho7rd...@news7.newsguy.com...
> SSTO, if it can be done at all with chemical fuels, is _barely_ doable.
There are several expendable stages which could theoretically do SSTO, with
a usable payload, if launched by themselves. Note that Atlas was able to
put Mercury into orbit, but it did cheat a bit by dropping the two outer
engines on the way up, partly to reduce thrust and partly to reduce the dry
mass of the booster.
That said, a resuable SSTO is a matter of debate. Some say it's possible,
others say it's too hard or impossible.
Jeff Findley
For a conventional bell engine design, quite a bit of ISP as you can
optimize the engine bell shape for vacuum.
J. Clarke
And that gains you what? Your SSTO still has to have enough fuel and
oxidizer aboard to impart 18,000 miles an hour of delta-v.
> at this point the airliner sets off its afterburners and releases the
> actual rocket stage, which achieves orbit.
The only airliners with afterburners are the Concorde and the TU-144,
neither of which under any circumstance can lift an X-33. In any case,
what do you believe that afterburners accomplish?
> the airliner flies back to base 100s if not a 1000 miles away.
And this gets you into orbit how? The question is not whether an
airliner can fly around with something attached to it, the question is
whether that thing that is attached can somehow achieve orbit. You have
not even attempted to address that question.
> a fully fuled rocket sitting on the pad is basically a loaded bomb.
>
> a airliner launched rocket stage can use ejection seats for the
> airliners crew, and a capsule safety pod for the rocket stage crew.
And what does this gain you? Does the crew jump out before the SSTO
actually starts for orbit or something? If your SSTO is to achieve
orbit at some point it has to become that "loaded bomb" that you fear
and if it is to take a crew into orbit then they have to be aboard that
"loaded bomb" that you fear, so how does attaching that loaded bomb to
an airliner change anything?
> think out of the box, the box isnt your friend..............
So show us the numbers that demonstrate that your "out of the box"
solution will work.
Oh, but that is that stupid boring math that is only for stupid boring
nerds and not for brilliant people like you, right?
J. Clarke
I'm asking the person making the assertion to back it up.
He is proposing that a kerosene-fueled X-33 attached to an airliner can
somehow achieve orbit.
I want to know how that is going to work.
What we know:
The X-33 was not designed to achieve orbit even with a hydrogen-oxygen
rocket.
<http://www.nasa.gov/centers/marshall/news/background/facts/x33.html>
Note maximum speed Mach 13. That is per
<http://www.aerospaceweb.org/design/scripts/atmosphere/> approximately
7500 knots true airspeed or about 8500 miles per hour. Orbital velocity
is approximately 18,000 miles/hr at 120 miles altitude per
<http://hyperphysics.phy-astr.gsu.edu/HBASE/orbv3.html>.
While I'm not going to give a cite for it, it is generally accepted that
all else being equal a kerosene rocket will have lower specific impulse
than a hydrogen rocket, so whatever performance the X-33 achieves with a
kerosene rocket will be less than for a hydrogen rocket.
So, tell me, how do you manage to get that additional 10,000 miles per
hour out of sticking the thing on top of an airliner?
Jeff Findley
> While I'm not going to give a cite for it, it is generally accepted that
> all else being equal a kerosene rocket will have lower specific impulse
> than a hydrogen rocket, so whatever performance the X-33 achieves with a
> kerosene rocket will be less than for a hydrogen rocket.
ISP is one measure of engine performance. Vehicle performance is much more
complicated and depends on many more variables besides engine ISP. In
particular, LH2 isn't very dense. Kerosene is far more dense than LH2 plus
it doesn't need cryogenic storage. In a vehicle design, kerosene has some
distinct advantages which may make up for its lower ISP.
J. Clarke
> "J. Clarke"<jclarke...@cox.net> wrote in message
> news:ho7rd...@news7.newsguy.com...
>> On 3/22/2010 8:32 AM, Jeff Findley wrote:
>>> "Greg D. Moore (Strider)"<mooregr_d...@greenms.com> wrote in
>>> message
>>> news:X6-dnTN9SfCfDzvW...@earthlink.com...
>>>> J. Clarke wrote:
>>>>> Why do people think that launching from 50,000 feet will help somehow?
>>>>> Going into orbit is not a matter of going high, it's a matter of going
>>>>> _fast_. Launching from 50,000 feet or from sea level you still need
>>>>> to impart 18,000 miles an hour of delta-v. That's the hard part.
>>>>
>>>> Because 50,000 feet gets you above the bulk of the atmosphere which
>>>> provides a decent bonus.
>>>
>>> Specifically, you can optimize your engines for the much lower pressure
>>> of
>>> 50,000 feet (to vacuum), as opposed to the compromises necessary to make
>>> them run at sea level.
>>
>> So how much do you think this gains you?
>
> For a conventional bell engine design, quite a bit of ISP as you can
> optimize the engine bell shape for vacuum.
How much Isp? And how much of the time during boost is it running in
vacuum?
> Jeff
J. Clarke
> "J. Clarke"<jclarke...@cox.net> wrote in message
> news:ho7rd...@news7.newsguy.com...
>> SSTO, if it can be done at all with chemical fuels, is _barely_ doable.
>
> There are several expendable stages which could theoretically do SSTO, with
> a usable payload, if launched by themselves.
Which would those be?
> Note that Atlas was able to
> put Mercury into orbit, but it did cheat a bit by dropping the two outer
> engines on the way up, partly to reduce thrust and partly to reduce the dry
> mass of the booster.
Yep, it's called a half-stage".
> That said, a resuable SSTO is a matter of debate. Some say it's possible,
> others say it's too hard or impossible.
In any case, do you think that it's going to be achieved by replacing
the hydrogen aerospike engines in the X-33 with something burning
kerosene and sticking it on top of an airliner?
J. Clarke
> news:ho8cv...@news6.newsguy.com...
>> While I'm not going to give a cite for it, it is generally accepted that
>> all else being equal a kerosene rocket will have lower specific impulse
>> than a hydrogen rocket, so whatever performance the X-33 achieves with a
>> kerosene rocket will be less than for a hydrogen rocket.
>
> ISP is one measure of engine performance. Vehicle performance is much more
> complicated and depends on many more variables besides engine ISP. In
> particular, LH2 isn't very dense. Kerosene is far more dense than LH2 plus
> it doesn't need cryogenic storage. In a vehicle design, kerosene has some
> distinct advantages which may make up for its lower ISP.
And those are going to put an X-33 in orbit?
Pat Flannery
> ISP is one measure of engine performance. Vehicle performance is much more
> complicated and depends on many more variables besides engine ISP. In
> particular, LH2 isn't very dense. Kerosene is far more dense than LH2 plus
> it doesn't need cryogenic storage. In a vehicle design, kerosene has some
> distinct advantages which may make up for its lower ISP.
Considering that it wasn't much larger than a V-2 and could put a small
satellite into polar orbit, the kerosene/hydrogen peroxide propellant
combo used on the first two stages of the British Black Arrow shouldn't
be overlooked either.
It certainly was an extremely clean-burning combo:
http://www.daviddarling.info/images/Black_Arrow.jpg
Making it look like the rocket was levitating more than lifting off.
I'm still shaking my head over the liquid fluorine/liquid deuterium
propellants proposed as one alternative for powering the ALSV.
I don't know what advantage liquid deuterium gives over stock liquid
hydrogen, but it had better be pretty impressive given what it's going
to cost to tank it up with that.
Pat
J. Clarke
Doubles the storage density.
Peter Stickney
It depends on the chamber pressure of the engine - but a fair bit -
the J2 engine optimized for Sea Level has a vacuum Isp of 390, and
the vacuum optimized J2 has an Isp of 421. = a gain of 8% over Sea Level.
A launch vehicle engine spends most of its time in vacuum. The initial
trajectory is as much vertical as possible to get it out of the thick air.
When a reasonably high altitude is reached, you pitch over to accelerate.
--
Pete Stickney
Failure is not an option
It comes bundled with the system
Peter Stickney
Pat - don't forget that he's responding to Bbo Hallrb, and therefore is arguing
with a twaddle of supreme ignorance.
Peter Stickney
While LH2 can provide high Isp, its Energy Density (Cubic Ergs, if you will)
is quite poor. Since an SSTO is fairly limited in volume, you need a high
energy density fuel.
Kerosene has about 6 times the energy density of LH2.
The drawback is it weighs more, and thus incurs structural weight penalties.
Jorge R. Frank
> On Mon, 22 Mar 2010 11:07:19 -0800, Pat Flannery wrote:
>
>> On 3/22/2010 6:54 AM, J. Clarke wrote:
>>> But your assertion does not convince. You are posting on the Internet.
>>> Most people posting on the Internet have opinions. Most of those
>>> opinions are ignorant twaddle. So one must take your opinion as
>>> ignorant twaddle until you can provide some numbers to go with it.
>> Yeah...but you are posting on the internet also, and I'm not seeing any
>> numbers so far. :-D
>
> Pat - don't forget that he's responding to Bbo Hallrb, and therefore is arguing
> with a twaddle of supreme ignorance.
>
Ah, yes. The Stimson J. Cat of Usenet himself.
J. Clarke
So you're saying that the Lockheed Skunk Works didn't know what they
were doing when they chose to use hydrogen?
J. Clarke
So what percentage of the time in a typical launch is spent in vacuum?
Peter Stickney
At least 80%
Pat Flannery
> "J. Clarke"<jclarke...@cox.net> wrote in message
> news:ho7rd...@news7.newsguy.com...
>> SSTO, if it can be done at all with chemical fuels, is _barely_ doable.
>
> There are several expendable stages which could theoretically do SSTO, with
> a usable payload, if launched by themselves.
I've never heard of one that could do that without dropping something on
the way up like Atlas did.
Someone here* suggested that Thor might be able to do it, but that
proved not to be the case.
* Someone who owns a lot of cats and a machine gun, IIRC. :-)
Pat
Robert Clark
> ... The
> advantages of high altitude fueling and launch for a Shuttle type
> vehicle to avoid ice buildup on tankage using any sort of cryogenic
> propellants go clean back to the Air Force/DARPA ALSV concept that
> Fluorine-deuterium? Oh, that will be cheap to use as fuel. ;-)
>
> Pat
Thanks for those links.
It has been rumored that the Air Force has tested such a system:
Two-Stage-to-Orbit ''Blackstar'' System Shelved at Groom Lake?
Mar 5, 2006
By William B. Scott
http://www.aviationweek.com/aw/generic/story_generic.jsp?channel=awst&id=news/030606p1.xml
TSTO spaceplanes.
http://robotpig.net/aerospace/en_tsto.php
Did Pentagon create orbital space plane?
Magazine reports evidence for classified project, sparking some
skepticism.
By James Oberg, NBC News space analyst
updated 5:10 p.m. ET, Mon., March. 6, 2006
http://www.msnbc.msn.com/id/11691989/
Bob Clark
Pat Flannery
Yeah, but look at the price for it in a gaseous state:
http://www.medicalisotopes.com/display_product.php?catnum=D1401&cat_id=105&alpha=~&caller=CATEGORY
Around a dollar a liter.
About the only time I heard of someone whipping up a lot of it in a
liquid form was for the "Mike" test of the prototype hydrogen bomb.
Pat
Robert Clark
> On 3/22/2010 9:55 AM, Jeff Findley wrote:
>
> >news:ho7rd...@news7.newsguy.com...
> >> SSTO, if it can be done at all with chemical fuels, is _barely_ doable.
>
> > There are several expendable stages which could theoretically do SSTO, with
> > a usable payload, if launched by themselves.
>
> I've never heard of one that could do that without dropping something on
> the way up like Atlas did.
> Someone here* suggested that Thor might be able to do it, but that
> proved not to be the case.
>
> * Someone who owns a lot of cats and a machine gun, IIRC. :-)
>
> Pat
The main example is the Titan II first stage:
Single-stage-to-orbit.
"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."
http://en.wikipedia.org/wiki/Single-stage-to-orbit#Dense_versus_hydrogen_fuels
This is notable since the Titan II was operational since the earliest
days of orbital rockets in the early 60's. The Titan II was being
fired even up to 2003 and there are still some left unfired. So we
could still do this proof of principle launch with a Titan II first
stage to prove SSTO is possible.
Such a launch with the Titan II or Falcon 1 first stages would be
fundamentally important to the future development of low cost space
access. It would have a similar effect as to the first breaking of the
sound barrier.
In the field of rocketry, it would be the single biggest advance
since Robert Goddard first successfully fired liquid-fueled rockets.
Bob Clark
Robert Clark
> On Mar 23, 3:38 am, Pat Flannery <flan...@daktel.com> wrote:
>
>
>
> > On 3/22/2010 9:55 AM, Jeff Findley wrote:
>
> > > "J. Clarke"<jclarke.use...@cox.net> wrote in message
> > >news:ho7rd...@news7.newsguy.com...
> > >> SSTO, if it can be done at all with chemical fuels, is _barely_ doable.
>
> > > There are several expendable stages which could theoretically do SSTO, with
> > > a usable payload, if launched by themselves.
>
> > I've never heard of one that could do that without dropping something on
> > the way up like Atlas did.
> > Someone here* suggested that Thor might be able to do it, but that
> > proved not to be the case.
>
> > * Someone who owns a lot of cats and a machine gun, IIRC. :-)
>
> > Pat
>
> The main example is the Titan II first stage:
>
> Single-stage-to-orbit.
> "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
>
> This is notable since the Titan II was operational since the earliest
> days of orbital rockets in the early 60's. The Titan II was being
> fired even up to 2003 and there are still some left unfired. So we
> could still do this proof of principle launch with a Titan II first
> stage to prove SSTO is possible.
> Such a launch with the Titan II or Falcon 1 first stages would be
> fundamentally important to the future development of low cost space
> access. It would have a similar effect as to the first breaking of the
> sound barrier.
> In the field of rocketry, it would be the single biggest advance
> since Robert Goddard first successfully fired liquid-fueled rockets.
>
See the discussion here for some other examples of SSTO's possible
since the earliest days of launch vehicles:
Newsgroups: sci.space.history
From: he...@spsystems.net (Henry Spencer)
Date: Sun, 21 Sep 2003 21:26:23 GMT
Subject: Re: Single stage to orbit, Atlas
http://groups.google.com/group/sci.space.history/msg/81cf0052339d7ec1?hl=en
Bob Clark
Robert Clark
> ...
>
> Note that the other half-scale suborbital demonstrators for the NASA
> RLV program by Rockwell and McDonnell-Douglas (see images linked
> below) could be built for comparable prices and would likewise become
> full orbital craft by switching to kerosene or other dense propellant.
> Then we could have 3 separate designs for fully reusable SSTO vehicles
> at costs that could allow fully private financing that would
> significantly reduce launch costs and would allow manned flights.
>
> Successful operation of these X-33-sized orbital vehicles at a profit
> would encourage private financing to build the full-scale VentureStar-
> sized RLV's that could bring launch costs down to the $100 to $200 per
> kilo range.
>
> Bob Clark
>
> http://www.astronautix.com/nails/x/x33rock.jpg
>
> http://www.astronautix.com/graphics/x/x33p4.jpg
For unknown reasons Google puts all sorts of extraneous junk at the
beginning and end of links in posts on Google Groups which prevented
those image links from operating. You can copy and paste the image
links in the address bar to pull up the images:
"http://www.astronautix.com/nails/x/x33rock.jpg"
"http://www.astronautix.com/graphics/x/x33p4.jpg"
Bob Clark