Re: A kerosene-fueled X-33 as a single stage to orbit vehicle.
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.
>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.
>...
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
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 – 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%, from 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 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 with the kerosene-engine
change, the payload would be 54,000, 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".
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
Robert Clark
[re-posted to correct typos.]
April 25 =96 27, 1996
Pressure vessel.
http://en.wikipedia.org/wiki/Pressure_vessel#Scaling
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".
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
Robert Clark
> for decades, it is still commonly believed that SSTO performance with
> chemical propulsion is not possible even among experts in the space
> industry ...
>
> 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*.
>
This last may seem a bit extreme. However, Burt Rutan in this recent
video interview about 8 minutes in noted that if rockets were operated
with the efficiencies of airliners, the price for a passenger to orbit
would be in the $12,000 range:
Big Think Interview With Burt Rutan.
A conversation with the aerospace engineer and founder of Scaled
Composites.
March 3, 2010
http://bigthink.com/ideas/18881
Robert Zubrin in his book, "Entering Space: Creating a Spacefaring
Civilization" comes to a similar conclusion:
"Current -day rockets, such as the kerosene/oxygen-fueled Atlas can
deliver about 1 percent of their takeoff mass to orbit - most (about
90 percent) of the remaining mass is propellant. The cost of a
kerosene/oxygen propellant mxture (at 3:1 oxygen/kerosene mixture
ratio) is about $0.20/kg. Since the propellant consumed during launch
has 90 times the mass of the payload delivered, the propellant cost of
sending a mass to orbit is about $18/kg. Assuming a total system
operating cost of six times the propellant cost (about double the
total cost/fuel ratio of airlines), the resulting price of a rocket
ride to orbit would be in the neighborhood of $100/kg, or $10,000 for
a 100-kg passenger. There is no fundamental reason why space-launch
prices in this range cannot be achieved."
"Entering Space", by Robert Zubrin, p.22-23.
Then it is notable that trans-Pacific flights are in this cost range.
For instance I earlier did a search on the Japan Airlines site for
round trip business class tickets from my town of Philadelphia to
Tokyo. It ranged from $6,600 to $21,000:
=========================================
Quote:
Select Your Flights
Philadelphia to Tokyo Thursday, April 9, 2009
Tokyo to Philadelphia Tuesday, April 14, 2009
Travelers: 1
Travel class: Business and First
Select your fare: Price differences within a fare type may be due to
flight connections or availability. Prices are per adult passenger and
include Taxes and Surcharges.
Fare type Fare description Lowest price
Business Saver Special Restricted. Bed-style seating on most long-haul
routes -
Executive Class. more details $6,672.48
Business Saver Restricted. Bed-style seating on most long-haul routes
-
Executive Class. more details $7,611.48
Business Normal Flexible. Bed-style seating on most long-haul routes -
Executive Class. more details $12,330.48
First Normal Flexible. World-renowned service and comfort - First
Class.
more details $21,589.48
=========================================
And first class tickets even one-way on Qantas from Los Angeles to
Australia are in this price range:
=========================================
Your Search
From: Los Angeles
To: Sydney
Depart:
Adults:
Children:
Infants:
Search Options:
Must travel on these dates
Flexible with dates
From: Los Angeles
To: Sydney
Depart Arrive Morning Afternoon Evening
Stops: Non-stop
All flights
Step 2 - Select your flight
Price displayed is for 1 adult and includes surcharges, fees and
taxes.
> View lowest fare around this date
Flights Out: Los Angeles to Sydney - Wed 24 Mar 10
From To Flight Business First
23:50 Los Angeles 08:25
(Fri) Sydney QF108 $7926 $15881
Duration: 14h 35m
22:30 Los Angeles 07:25
(Fri) Sydney QF12 $7926 $15881
Duration: 14h 55m
=========================================
Then recall for the full-sized VentureStar RLV by switching to
kerosene we could get in the range of $166/kg launch costs or $16,600
for a 100 kg passenger.
Bob Clark
Jeff Findley
> On 3/22/2010 1:55 PM, Jeff Findley wrote:
>> "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?
Google is your friend. Try Googling for something like "expendable ssto
henry spencer". It's been discussed here several times, when this group had
a far better signal to noise ratio. Unfortunately, Henry Spencer no longer
posts here. Sigh...
>> 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".
But as Henry Spencer said of Atlas in one of his posts:
More precisely, its first stage could have taken 1-2klb of payload into
orbit all by itself, assuming suitable engines with the same Isp and
engine mass as the standard ones.
In other words, the stage as flown couldn't be considered an SSTO, but a bit
of development to produce a deep throttling sustainer engine could have made
it an SSTO. Atlas easily had the mass fraction and engine performance
necessary, but it lacked the deep throttling necessary to make it happen.
>> 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?
I'm not going to argue that point. But you might want to Google Black
Horse. It's a very interesting concept on which there was extensive number
crunching done. Oh heck, it's interesting enough I'll GIVE you a link:
http://www.ai.mit.edu/projects/im/magnus/bh/analog.html
This is the sort of "outside the box" thinking for a (near) SSTO that
doesn't violate physics and relies on the proven technology of in flight
refueling, which is done routinely by the military. Heck, even Air Force
One is equipped to receive fuel via this method. If it's safe enough and
reliable enough for the President of the United State's aircraft, it ought
to be good enough for a launch system!
Jeff Findley
You're the one claiming the virtues of a high ISP fuel like LH2/LOX, so I
assumed you were one of those "performance uber alles" types who would know
these things. Some engines have sea level and vacuum versions. Do a little
research and find the data for yourself. Either that or do the math (i.e.
college level aerospace engineering stuff).
To partially get around this very issue, the SSME's run at a relatively high
chamber pressure compared to other engines. These high chamber pressures
create issues of their own. The SSME has literally taken decades to
perfect.
Jeff Findley
I never claimed they would, I claim that LH2 isn't necessarily the best
choice for a rocket fuel because all else is NOT equal when you change the
fuel in a vehicle design. You can't look at a rocket engine's specs and
immediately conclude that LH2 is superior to kerosene. The devil is in the
details, particularly the details of the *vehicle* design.
As an example, compare the Saturn V first stage design to the second stage
design. Note the relative difference in size of the kerosene/LOX tanks on
the first stage to the LH2/LOX tanks on the second stage. The difference is
quite striking. Then as a thought experiment, try to figure out how big the
Saturn V first stage would need to have been if the fuel was LH2 instead of
kerosene.
Jeff Findley
Perhaps they knew their customer has much experience with LH2 and believes
that LH2 is the "best" fuel to use. Think about it. Lockheed got paid
regardless of the project success. Their goal was to win the contract,
which means give the customer what the customer thinks they want, not what
will actually succeed.
In the end, the X-33 failed, but Lockheed got paid and their existing EELV
related contracts continued. So from their point of view (i.e. upper
management), did they really fail from a business point of view?
Jeff Findley
For someone who was asking others to "do the math", you seem incapable of
that task yourself. Furthermore, you seem incapable of using Google to
look for information from others who have done the math. What's up with
that?
Jeff Findley
"Pat Flannery" <fla...@daktel.com> wrote in message
news:5audncWlgt8o2DXW...@posted.northdakotatelephone...
Henry Spencer did the math for several existing rocket stages. What most
needed to make this happen would be a deeply throttlable engine. Atlas was
an example. From memory, one of the Titan II stages and I think one of the
Saturn V stages also had the appropriate mass fraction.
If you do the math an expendable SSTO isn't really that hard.
Pat Flannery
>> 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.
>
> So you're saying that the Lockheed Skunk Works didn't know what they
> were doing when they chose to use hydrogen?
VentureStar relied on its large size when its propellants were exhausted
to reduce reentry heating to the point where fragile Shuttle-type tiles
weren't going to be needed and more robust metallic tiles could be
substituted for them.
That having been said, LH2 had once before led Lockheed astray:
http://en.wikipedia.org/wiki/Lockheed_CL-400_Suntan
In retrospect, the metallic tile concept for VentureStar may have been a
flop if employed, as they were supposed to be based on titanium, and as
was discovered after Columbia broke up on reentry, titanium burns at a
lower temperature than aluminum when in a atomic oxygen-rich
environment; so the first orbital flight of the VentureStar prototype
would have probably been its last.
Assuming they could have fixed that somehow, VentureStar was only
missing its performance goals marginally, due to the need to replace the
composite LH2 tank with a aluminum-lithium one, and the added weight and
drag of the twin vertical fins that were found to be necessary to assure
the vehicle's stability during atmospheric flight to and from orbit.
It would have taken a monster aircraft to carry it, but air-launching
the VentureStar at high altitude might have given it enough added "umph"
to have made it work as designed.
Pat
J. Clarke
What's up with that is that if you make the assertion it's up to you to
defend it.
If you think that a kerosene fueled X-33 dropped off an airliner is a
good idea then say why. If you think it isn't, say why. If you have no
opinion on it and want to discuss some unrelated system, please start a
new thread.
J. Clarke
The X-33 is not VentureStar, VentureStar is not the X-33, the X-33 was a
subscale prototype never intended to achieve orbit. So why are you
bringing VentureStar into the discussion?
J. Clarke
> "J. Clarke"<jclarke...@cox.net> wrote in message
> news:ho8gg...@news5.newsguy.com...
>> On 3/22/2010 2:37 PM, Jeff Findley wrote:
>>> "J. Clarke"<jclarke...@cox.net> wrote in message
>>> 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?
>
> I never claimed they would,
Then why are you introducing them to this thread?
<Remainder, with no relevance to the kerosene-fueled X-33 concept snipped>
J. Clarke
Black Horse is an interesting concept but much different from the notion
of sticking some kerosene fueled engines in the X-33 design and dropping
it off a commercial airliner.
>
> Jeff
J. Clarke
> "Pat Flannery"<fla...@daktel.com> wrote in message
> news:5audncWlgt8o2DXW...@posted.northdakotatelephone...
>> On 3/22/2010 9:55 AM, Jeff Findley wrote:
>>> "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. :-)
>
> Henry Spencer did the math for several existing rocket stages. What most
> needed to make this happen would be a deeply throttlable engine. Atlas was
> an example. From memory, one of the Titan II stages and I think one of the
> Saturn V stages also had the appropriate mass fraction.
>
> If you do the math an expendable SSTO isn't really that hard.
So why are they not in common use?
J. Clarke
I see, so the Lockheed Skunk Works is in the business of designing
flying machines that can't fly?
J. Clarke
> "J. Clarke"<jclarke...@cox.net> wrote in message
> news:ho8eo...@news5.newsguy.com...
>> On 3/22/2010 1:56 PM, Jeff Findley wrote:
>>> 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?
>
> You're the one claiming the virtues of a high ISP fuel like LH2/LOX, so I
> assumed you were one of those "performance uber alles" types who would know
> these things. Some engines have sea level and vacuum versions. Do a little
> research and find the data for yourself. Either that or do the math (i.e.
> college level aerospace engineering stuff).
>
> To partially get around this very issue, the SSME's run at a relatively high
> chamber pressure compared to other engines. These high chamber pressures
> create issues of their own. The SSME has literally taken decades to
> perfect.
So you're saying that the space shuttle program ran its entire course
from development to retirement without a "perfected" engine?
And what does this have to do with a kerosene fueled X-33?
Jeff Findley
So you want to be able to make sweeping generalizations well beyond the
context of the subject line but sill be able to hide behind the subject line
when those same sweeping generalizations are shot down? Wow.
Pat Flannery
> But as Henry Spencer said of Atlas in one of his posts:
>
> More precisely, its first stage could have taken 1-2klb of payload into
> orbit all by itself, assuming suitable engines with the same Isp and
> engine mass as the standard ones.
>
> In other words, the stage as flown couldn't be considered an SSTO, but a bit
> of development to produce a deep throttling sustainer engine could have made
> it an SSTO. Atlas easily had the mass fraction and engine performance
> necessary, but it lacked the deep throttling necessary to make it happen.
Originally, the Atlas A test version (as the X-11) was just going to use
the central sustainer engine, rather than using just the two outer
booster engines as was actually flown.
> I'm not going to argue that point. But you might want to Google Black
> Horse. It's a very interesting concept on which there was extensive number
> crunching done. Oh heck, it's interesting enough I'll GIVE you a link:
>
> http://www.ai.mit.edu/projects/im/magnus/bh/analog.html
Robert Zubrin thinks this is a good idea that would be easy to
implement; that should warn you right there.
> This is the sort of "outside the box" thinking for a (near) SSTO that
> doesn't violate physics and relies on the proven technology of in flight
> refueling, which is done routinely by the military. Heck, even Air Force
> One is equipped to receive fuel via this method. If it's safe enough and
> reliable enough for the President of the United State's aircraft, it ought
> to be good enough for a launch system!
The difference being in the case of Black Horse what it was being tanked
up with at altitude was hydrogen peroxide, not jet fuel.
Rather than going the aerial refueling route, it would be a lot easier
to carry the fully-fueled Black Horse to altitude atop a 747 or C-5B,
and air-launch it, or air launch it after fueling it in-flight from
internal supplies while still attached to the carrier aircraft.
You can then make the landing gear even lighter, as it's now only used
for landing with no propellants aboard.
Once you do that, Black Horse suddenly becomes very similar to the Air
Launched Sortie Vehicle in concept, though strangely needing far less of
a propellant load to reach orbit - ALSV needed one or more drop tanks to
accomplish this.
Take a look at the weight estimates in that article versus the amount of
propellant it's supposed to carry; empty weight is 14,958 lbs, ignition
for the climb to orbit weight is 184,250 lbs - a mass ratio of 12.3.
Atlas had a mass ratio identical to that:
http://www.brighthub.com/science/space/articles/29607.aspx
...despite lacking goodies like wings, landing gear, a cockpit, added
jet engines, and a TPS...and possessing a shape that was far more
efficient for storing propellants in for external area. Even then, it
needed to jettison a couple of its engines to make it into orbit.
The H2O2/JP-5 propellant combo for Black Horse also has lower isp than
the Atlas LOX/kerosene. LOX is pure oxygen rather than a oxygen/hydrogen
mixture so you are also getting superior reactive energy storage with it
per volume over H2O2, despite the higher density of H2O2 (1.463 g/cm3
vs. 1.141 g/cm3 for LOX).
So how is this all supposed to work?
The answer is that it doesn't work; the whole thing is a bunch of hooey
thrown together out of badly unrealistic mass and performance estimates.
Pat
J. Clarke
Again comments with nothing to do with the proposed concept. I bet
you're real fun in design reviews.
Pat Flannery
>
> Henry Spencer did the math for several existing rocket stages. What most
> needed to make this happen would be a deeply throttlable engine. Atlas was
> an example. From memory, one of the Titan II stages and I think one of the
> Saturn V stages also had the appropriate mass fraction.
Now that you mention it, I seem to remember statements that the Titan II
first stage could do this, but with almost no payload aboard.
This claims that for it: http://en.wikipedia.org/wiki/Single-stage-to-orbit
But also states that it has a mass fraction of 25/1; according to
Encyclopedia Astronautica the stage had a mass fraction of 17.5/1:
http://www.astronautix.com/lvs/titan.htm
From here, it's 23.7/1: http://www.titan2icbm.org/titanD.html
That seems awfully high, even given the chem-milled tankage.
There was also a proposal for a launcher based on the first stage of the
Saturn V, but that one was to jettison the four outer F-1s on the way
up: http://www.up-ship.com/drawndoc/sdoc63ad.jpg
> If you do the math an expendable SSTO isn't really that hard.
Yeah, but the payload is certainly nothing to get excited about, and
that right there should have been a warning about reusable ones that
need TPS and recovery systems.
Pat
Pat Flannery
>> Henry Spencer did the math for several existing rocket stages. What most
>> needed to make this happen would be a deeply throttlable engine. Atlas
>> was
>> an example. From memory, one of the Titan II stages and I think one of
>> the
>> Saturn V stages also had the appropriate mass fraction.
>>
>> If you do the math an expendable SSTO isn't really that hard.
>
> So why are they not in common use?
Because of the low payload that they can carry into orbit versus the
overall cost of the vehicle.
If adding a second stage increases vehicle cost by 75% but allows you to
put twice as much weight into orbit, then your price-per-pound into
orbit is lower than using a SSTO launcher.
Pat
Pat Flannery
> The X-33 is not VentureStar, VentureStar is not the X-33, the X-33 was a
> subscale prototype never intended to achieve orbit. So why are you
> bringing VentureStar into the discussion?
Because the two programs were related.
Robert Clarke keeps trying to come up with some way to get the X-33 into
orbit by fiddling with its design in some way.
So I'm going to fiddle with it too.
I'm going to make it a hell of a lot bigger and call it VentureStar. ;-)
Pat
Pat Flannery
>> In the end, the X-33 failed, but Lockheed got paid and their existing
>> EELV
>> related contracts continued. So from their point of view (i.e. upper
>> management), did they really fail from a business point of view?
>
> I see, so the Lockheed Skunk Works is in the business of designing
> flying machines that can't fly?
They've done that before with the CL-400 project.
Sometimes they bite off more than they can chew, and the X-33 was a case
in point. Considering all the trouble the A-12 gave them, that could
have gone the other way too...and the F-35 isn't being a pushover to
build either.
Pat
Pat Flannery
> So you want to be able to make sweeping generalizations well beyond the
> context of the subject line but sill be able to hide behind the subject line
> when those same sweeping generalizations are shot down? Wow.
Watch it mate.
E's the bleedin' Pope now. :-D
Pat
J. Clarke
Well that's fine but it's not what he's proposing.
My point in all this is that the only way he's going to get X-33 into
orbit is to stick it on top of a Saturn V and that his notion of
sticking it on an airliner buys him nothing.
And unless he happens to have the remains sitting in his barn I don't
understand why he's so hot to recycle the X-33 anyway.
J. Clarke
> On 3/23/2010 6:53 AM, J. Clarke wrote:
>>> In the end, the X-33 failed, but Lockheed got paid and their existing
>>> EELV
>>> related contracts continued. So from their point of view (i.e. upper
>>> management), did they really fail from a business point of view?
>>
>> I see, so the Lockheed Skunk Works is in the business of designing
>> flying machines that can't fly?
>
> They've done that before with the CL-400 project.
In what way was that designed to not fly? Pratt promised them an
engine. I've worked for UTC and it my experience there was such that it
does not surprise me that Pratt failed to deliver. But if Pratt had
delivered the engine would it still not have flown and why not?
> Sometimes they bite off more than they can chew, and the X-33 was a case
> in point.
X-33 was an X-plane whose purpose was to find out what problems would
come about in developing VentureStar. They found the problems. The
customer decided not to fix them. But it's not because X-33 was
designed to not fly.
> Considering all the trouble the A-12 gave them, that could
> have gone the other way too...
But it didn't because the customer decided to finish the project for once.
> and the F-35 isn't being a pushover to
> build either.
If you are using that to back an assertion that the Skunk Works designs
planes to not fly, you have chosen a singularly poor example.
You seem to be equating "had development problems" with "designed to not
fly".
J. Clarke
In other words they can barely struggle into orbit rather than being
"not that hard".
Michal Jankowski
>>> If you do the math an expendable SSTO isn't really that hard.
>>
>> So why are they not in common use?
>
> Because of the low payload that they can carry into orbit versus the
> overall cost of the vehicle.
...because nobody even tries to build them, and that's mostly because
nobody's ever built any.
http://sci.tech-archive.net/Archive/sci.space.history/2007-07/msg00146.html
MJ
Marvin the Martian
> On 3/21/2010 4:53 PM, Marvin the Martian wrote:
>
>
>> 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
If pigs could fly, then you could tie an apple to your rocket and have
them take your rocket to space.
Yeah, right. You're going to dock with a refueling rocket to send up the
oxidizer on an accelerating rocket for the brief moments that it is on
it's way up.
That problem has been addressed by staging the rocket. The "refuel"
rocket is the first stage.
Now get real.
Marvin the Martian
> On Mar 22, 10:54�am, "J. Clarke" <jclarke.use...@cox.net> wrote:
>> On 3/22/2010 10:20 AM, hall...@aol.com wrote:
>>
>>
>>
>>
>>
>> > On Mar 22, 9:12 am, "J. Clarke"<jclarke.use...@cox.net> �wrote:
>> >> On 3/22/2010 4:07 AM, Pat Flannery wrote:
>>
>> >>> On 3/21/2010 4:53 PM, Marvin the Martian wrote:
>>
>> >>>> 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.
>>
>> >> 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.- Hide quoted text -
>>
>> >> - 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.
>>
>> 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.- Hide quoted text -
>>
>> - 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..............
You're an idiot, spouting off about something you're utterly ignorant and
not even bothering to check out the physics. I'm all for THINKING out of
the box, but totally against drug induced delusional monkey gibber.
To get to orbit, your minimum speed is 7 km/sec. The fastest airplanes
can travel about .9 km/sec, and those are fighter jets, not tankers.
You're talking about an energy difference of 7^-0.9^2 = 49 - 0.8, or
0.8/49 =~ 2%. And for that piddly advantage, you have to including
docking, refueling hardware, and screw around with transferring
CRYOGENIC liquid oxygen at super sonic speeds.
That's just stupid on the face of it.
Marvin the Martian
> "J. Clarke" <jclarke...@cox.net> wrote in message
> 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.
>
> Jeff
The paper I read on kerosene rockets argued that the rocket mass was
increased so much because of the size of the liquid hydrogen tank and
compressors to keep it liquid such that you could actually loft more
payload (but less total rocket mass) with a kerosene rocket than a
hydrogen one, since kerosene requires no pumps, no insulation, a smaller
tank, and far less support structure than liquid hydrogen does.
The paper made sense.
This thread doesn't make sense. It is just absurd statements put out by
some gibbering wannabe posers. The same arguments for staged rockets
rather than single stage to orbit still stands. The gibber about
transferring LOX in the boost phase is just stupid unthinking gibber.
Pat Flannery
>> 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
>
> If pigs could fly, then you could tie an apple to your rocket and have
> them take your rocket to space.
>
> Yeah, right. You're going to dock with a refueling rocket to send up the
> oxidizer on an accelerating rocket for the brief moments that it is on
> it's way up.
Nothing like that at all; any cryogenic propellants are stored in
insulated tanks aboard the carrier aircraft and transferred via pipes
into the orbital spacecraft (or its drop tanks) riding atop or under it
once the carrier aircraft is at altitude, and temperature and liquid
atmospheric moisture are very low.
Once it's time to launch, the orbital spacecraft detaches and fires its
engines.
That was the concept they were going to use on the 747-launched Air
Force minishuttle of the 1980's to eliminate the need for foam
insulation on its big drop tank:
http://danielmarin.blogspot.com/2008/12/maks-y-alsv.html
This Rockwell study would appear to carry the two-stage vehicle inside
of a humidity controlled interior of a C-5B to accomplish the same
prevention of ice build-up: http://www.thespacereview.com/archive/1591a.jpg
Pat
Pat Flannery
>
> To get to orbit, your minimum speed is 7 km/sec. The fastest airplanes
> can travel about .9 km/sec, and those are fighter jets, not tankers.
>
> You're talking about an energy difference of 7^-0.9^2 = 49 - 0.8, or
> 0.8/49 =~ 2%. And for that piddly advantage, you have to including
> docking, refueling hardware, and screw around with transferring
> CRYOGENIC liquid oxygen at super sonic speeds.
>
> That's just stupid on the face of it.
If it's a subsonic aircraft, you only get the advantage of lower ascent
air drag by launching in the thin air at altitude (which is why the
sub-orbital SpaceShip 2 will get launched at altitude from its carrier
aircraft rather than being verticaly launched from the ground.)
But as the capabilities of the carrier aircraft increase, it can be
turned into the equvelent of a reusable first stage for a multi-stage
launch system.
In the Soviet "Spiral" design, the carrier aircraft was to get up to
Mach 6-7 before releasing the spaceplane riding on the nose of its own
rocket second stage: http://www.buran.ru/htm/str126.htm
pat
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.
>>
>> Jeff
>
> The paper I read on kerosene rockets argued that the rocket mass was
> increased so much because of the size of the liquid hydrogen tank and
> compressors to keep it liquid
No refrigeration compressors are used on LH2 fueled boosters.
The LH2 is put aboard and topped up to make up for boil-off till just
shortly before launch.
Pat
J. Clarke
You haven't shown that insulation of the tank is the major obstacle to
SSTO. As has been pointed out, it didn't stop Atlas from coming very
close to achieving that objective with '50s material technology.
J. Clarke
> On 3/24/2010 6:33 PM, Marvin the Martian wrote:
>>
>> To get to orbit, your minimum speed is 7 km/sec. The fastest airplanes
>> can travel about .9 km/sec, and those are fighter jets, not tankers.
>>
>> You're talking about an energy difference of 7^-0.9^2 = 49 - 0.8, or
>> 0.8/49 =~ 2%. And for that piddly advantage, you have to including
>> docking, refueling hardware, and screw around with transferring
>> CRYOGENIC liquid oxygen at super sonic speeds.
>>
>> That's just stupid on the face of it.
>
> If it's a subsonic aircraft, you only get the advantage of lower ascent
> air drag by launching in the thin air at altitude (which is why the
> sub-orbital SpaceShip 2 will get launched at altitude from its carrier
> aircraft rather than being verticaly launched from the ground.)
It's not trying to go fast, it's trying to go high. For that every
little bit helps.
> But as the capabilities of the carrier aircraft increase, it can be
> turned into the equvelent of a reusable first stage for a multi-stage
> launch system.
> In the Soviet "Spiral" design, the carrier aircraft was to get up to
> Mach 6-7 before releasing the spaceplane riding on the nose of its own
> rocket second stage: http://www.buran.ru/htm/str126.htm
That's a bit different from sticking an X-33 on a commercial airliner.
I don't think that you'll find many people arguing that a design using a
Mach 7 carrier aircraft is undesirable, IF we can build the Mach 7
carrier aircraft, which nobody has ever succeeded in doing. Show us
something big sustaining Mach 7 in level powered flight.
Jeff Findley
True. That's why it's called a cryogenic fuel. It's liquid mostly because
it's been chilled to cryogenic temperatures, not because it has been
compressed to very high pressure.
Pat Flannery
> You haven't shown that insulation of the tank is the major obstacle to
> SSTO. As has been pointed out, it didn't stop Atlas from coming very
> close to achieving that objective with '50s material technology.
It certainly would be possible to make a expendable SSTO with today's
technology...they almost got VentureStar to work, and that had to return
to Earth after placing its payload in orbit.
The problem is an economic one as far as expendable SSTO payload
fraction versus overall launch cost.
As I pointed out earlier, a rocket possessing a second stage can more
than double the payload placed in orbit without doubling the launch
cost, which is why no one was really interested in building a SSTO
expendable system.
Discussion has been made here about sticking deep throttling engines on
the Atlas to give it true SSTO ability without having to drop the
booster engines during ascent.
While that may have been possible to do, does it give any real advantage
over the way it was actually done?
You still lose the whole vehicle and its three engines on every mission,
and the deep throttling engines probably cost more than the ones it
actually used.
So other than proving a philosophical point, the SSTO version doesn't
give you any real advantage.
The SSTO Titan II is an interesting concept (assuming it really could do
that - that 25/1 mass fraction seems off), but one has to ask oneself if
you could use a smaller two-stage rocket of lower cost to orbit the same
usable payload weight as the Titan II's first stage?
Note that the descriptions of the SSTO Titan II don't even talk about it
carrying a payload, just putting itself into orbit.
The Saturn V first stage with the droppable outboard engines might have
made some sense if you could have figured out some way to recover the
four outboard F-1s for re-use, or maybe just as it was designed, as a
expendable super-Atlas concept.
We'll never know on that one, as all the advanced Saturn V concepts were
killed by the Shuttle program, if for no other reason than to allow the
Saturn V launch pads to be converted for Shuttle use. If the Shuttle had
met the rosy economic predictions that were originally claimed for it -
with all the SRB segments being re-used and the orbiter being very easy
to refurbish for relaunch - it might have been even cheaper to fly than
the expendable Saturn V first stage vehicle.
Pat
Marvin the Martian
> On 3/24/2010 6:33 PM, Marvin the Martian wrote:
>>
>> To get to orbit, your minimum speed is 7 km/sec. The fastest airplanes
>> can travel about .9 km/sec, and those are fighter jets, not tankers.
>>
>> You're talking about an energy difference of 7^-0.9^2 = 49 - 0.8, or
>> 0.8/49 =~ 2%. And for that piddly advantage, you have to including
>> docking, refueling hardware, and screw around with transferring
>> CRYOGENIC liquid oxygen at super sonic speeds.
>>
>> That's just stupid on the face of it.
>
> If it's a subsonic aircraft, you only get the advantage of lower ascent
> air drag by launching in the thin air at altitude (which is why the
> sub-orbital SpaceShip 2 will get launched at altitude from its carrier
> aircraft rather than being verticaly launched from the ground.)
"Space ship 2" is a damned sounding rocket that amazes the simple minded.
IT doesn't get anywhere NEAR orbital velocity, so yeah, carrying it aloft
makes sense.
That's a big apples and oranges comparison.
< snip blather >
What the hell was that about? I just explained why it wouldn't work.
Don't be obtuse.
Marvin the Martian
> On 3/24/2010 6:24 PM, Marvin the Martian wrote:
>>> 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
>>
>> If pigs could fly, then you could tie an apple to your rocket and have
>> them take your rocket to space.
>>
>> Yeah, right. You're going to dock with a refueling rocket to send up
>> the oxidizer on an accelerating rocket for the brief moments that it is
>> on it's way up.
>
> Nothing like that at all; any cryogenic propellants are stored in
> insulated tanks aboard the carrier aircraft and transferred via pipes
> into the orbital spacecraft (or its drop tanks) riding atop or under it
> once the carrier aircraft is at altitude, and temperature and liquid
> atmospheric moisture are very low.
Like I said, that's stupid for the reasons given. Repeating the same
stupidity over and over again is NOT an argument.
> Once it's time to launch, the orbital spacecraft detaches and fires its
> engines.
> That was the concept they were going to use on the 747-launched Air
> Force minishuttle of the 1980's to eliminate the need for foam
> insulation on its big drop tank:
> http://danielmarin.blogspot.com/2008/12/maks-y-alsv.html This Rockwell
> study would appear to carry the two-stage vehicle inside of a humidity
> controlled interior of a C-5B to accomplish the same prevention of ice
> build-up: http://www.thespacereview.com/archive/1591a.jpg
And it was never done because it's STUPID.
David Spain
> rather than single stage to orbit still stands. The gibber about
> transferring LOX in the boost phase is just stupid unthinking gibber.
Well not transferring from an external tanker, but consumption from
the carrier aircraft in a TSTO vehicle was seriously considered
by Boeing.
I've taken a look at the US Patent (4,802,639) assigned to Boeing in
1989 for the 'Horizontal-Takeoff Transatmospheric Launch System'
and in its 'first embodiment' the booster vehicle and the orbiter
vehicle share the same LH2/LO2 tankage using two SSME's one on the
booster and one on the orbiter. Some other relevant excerpts from the
patent application:
Page 14 starting at section 60:
'In addition to the structural connection between the aircraft 2 and
vehicle 50 provided by the struts 30,32, before separation the fuel
systems of the aircraft 2 and vehicle 50 are interconected to provide
cross-feeding of rocket fuel from the aircraft 2 to the vehicle 50.
This cross-feeding ensures that the vehicle 50 has essentially full fuel
tanks and carries a maximum amount of rocket fuel when the two stages 2,
50 separate. Fig. 17 is a schematic view of the rocket fuel systems of
the aircraft 2 and vehicle 50 and the corss-feeding means. As can be
seen in Fig. 17, a fuel line 25 from the aircraft liquid hydrogen tanks
24 communicates with a conduit in a couply 88 for transfer to the vehicle
50. This coupling 88 is located on an exterior surface in the cavity 4.
A shut-off valve 86 is provided to seal off the line 25 when the separation
procedure commences. Similarly, a fuel line 28 from the liquid oxygen tanks
26 communicates with a separate conduit in the coupling 88 for transfer
to the vehicle 50. Line 28 has a shut-ff valve 82. The coupling 88 on the
aircraft 2 mates with a coupling 90 oon an exterior surface of the vehicle
50. Together, couplings 88, 90 form a releasable connection which may be
of various known types, such as a poppet valve connection with restraints.
Conduits 96, 98 from the coupling 90 feed the fuel to the main rocket
engine 66 of the vehicle 50. These conduits are provided with shut-off
valves 92, 94 to close off the coupling 90 when the vehicle 50 is separated
from the aircraft 2.'
The numbers refer to elements shown in Figure 17 of the patent.
'Aircraft 2' refers to the carrier craft, 'vehicle 50' is the orbiter.
In Fig. 10 of the patent is a proposed flight tragectory for both the
booster aircraft and the orbiter. At an altitude of 30,000 ft. and
Mach number of .85, the SSMEs on both carrier and orbiter are ignited
and flown up to the separation point at 103,800 ft and Mach 3.3.
At this point the fuel systems are disconnected, the orbiter switches
to its own internal fuel supply and continue to orbit. The Booster
throttles its SSME down during the release procedure and levels off
to a much shallower ascent curve maxing at an altitude of 127,000 ft
and a reduced Mach number of 2.84 and then flies down from there
back to base at some point switching back to its 8 air-breathing jet
engines.
What is left out in the 'first embodiment' proposal in this patent
is how you modify an SSME to be self-starting. I'm speculating the
most reliable method might be using hypergolic pre-start injectors.
You'd need to add these to both the SSMEs as well as to the LH2/LOX
OMS thrusters as proposed for 'vehicle 50' in this patent. Any work
done in this area that has been documented?
BTW, how much experience is there when dealing with hybrid aircraft
that have both rocket propulsion and air-breathing engines? At the
proposed altitudes published for the booster aircraft, I would
think there are some significant challenges there in switching between
an airbreathing jet engine and a rocket and back again...
Any links to known studies appreciated...
Dave
Pat Flannery
>> That was the concept they were going to use on the 747-launched Air
>> Force minishuttle of the 1980's to eliminate the need for foam
>> insulation on its big drop tank:
>> http://danielmarin.blogspot.com/2008/12/maks-y-alsv.html This Rockwell
>> study would appear to carry the two-stage vehicle inside of a humidity
>> controlled interior of a C-5B to accomplish the same prevention of ice
>> build-up: http://www.thespacereview.com/archive/1591a.jpg
>
> And it was never done because it's STUPID.
If it was stupid, it's a stupid that both the US and Russia put a lot of
effort into studying over the years.
One advantage you get is you can fly out over the ocean and launch into
any orbital inclination you want without worrying about stages coming
down in inhabited areas.
Pat
Pat Flannery
> Marvin the Martian<mar...@ontomars.org> writes:
>> rather than single stage to orbit still stands. The gibber about
>> transferring LOX in the boost phase is just stupid unthinking gibber.
>
> Well not transferring from an external tanker, but consumption from
> the carrier aircraft in a TSTO vehicle was seriously considered
> by Boeing.
They are doing a multi-part series on that concept over on The Space Review:
http://www.thespacereview.com/article/1569/1
http://www.thespacereview.com/article/1580/1
http://www.thespacereview.com/article/1591/1
With the Boeing study being the next part scheduled to appear.
> What is left out in the 'first embodiment' proposal in this patent
> is how you modify an SSME to be self-starting. I'm speculating the
> most reliable method might be using hypergolic pre-start injectors.
> You'd need to add these to both the SSMEs as well as to the LH2/LOX
> OMS thrusters as proposed for 'vehicle 50' in this patent. Any work
> done in this area that has been documented?
Would you believe spark plugs on the J-2?:
http://www.apollosaturn.com/s5news/p61-7.htm
"AUGMENTED SPARK IGNITER
"The augmented spark igniter (ASI) is mounted to the injector face. It
provides the flame to ignite the propellants in the thrust chamber. Then
engine start is initiated, the spark exciters energize two spark plugs
mounted in the side of the igniter chamber. Simultaneously, the control
system starts the initial flow of oxidizer and fuel to the spark
igniter. As the oxidizer and fuel enter the combustion chamber of the
ASI, they mix and are ignited.
Mounted in the ASI is an ignition monitor which indicates that proper
ignition has taken place. The ASI operates continuously during entire
engine firing, is uncooled, and is capable of multiple reignitions under
all environmental conditions."
The shuttle uses the same basic ignition system:
http://spaceflight.nasa.gov/shuttle/reference/shutref/orbiter/prop/engines.html
"The oxidizer and fuel preburners are welded to the hot-gas manifold.
The fuel and oxidizer enter the preburners and are mixed so that
efficient combustion can occur. The augmented spark igniter is a small
combination chamber located in the center of the injector of each
preburner. The two dual-redundant spark igniters, which are activated by
the engine controller, are used during the engine start sequence to
initiate combustion in each preburner. They are turned off after
approximately three seconds because the combustion process is then
self-sustaining. The preburners produce the fuel-rich hot gas that
passes through the turbines to generate the power to operate the
high-pressure turbopumps. The oxidizer preburner's outflow drives a
turbine that is connected to the HPOT and the oxidizer preburner pump.
The fuel preburner's outflow drives a turbine that is connected to the
HPFT."
Pat
Pat Flannery
>> rather than single stage to orbit still stands. The gibber about
>> transferring LOX in the boost phase is just stupid unthinking gibber.
>
> Well not transferring from an external tanker, but consumption from
> the carrier aircraft in a TSTO vehicle was seriously considered
> by Boeing.
>
> I've taken a look at the US Patent (4,802,639) assigned to Boeing in
> 1989 for the 'Horizontal-Takeoff Transatmospheric Launch System'
> and in its 'first embodiment' the booster vehicle and the orbiter
> vehicle share the same LH2/LO2 tankage using two SSME's one on the
> booster and one on the orbiter. Some other relevant excerpts from the
> patent application:
Two things about this patent:
1.) Google now has a patent search engine out in a Beta version that let
me find and download a PDF of it: http://www.google.com/patents
2.) Some consider this to be related to the "Blackstar" TSTO vehicle
AW&ST had the cover story about:
http://www.aviationweek.com/aw/generic/story_generic.jsp?channel=awst&id=news/030606p1.xml
...being Boeing's entry in a competition to develop a system like this
that was won by another aerospace firm (Lockheed or Rockwell).
Pat
Pat Flannery
> Two things about this patent:
> 1.) Google now has a patent search engine out in a Beta version that let
> me find and download a PDF of it: http://www.google.com/patents
Entering "4,802,639" into that search engine will bring up a whole five
pages worth of related vehicle patents, most of which I've never seen
before.
Pat
David Spain
> Two things about this patent:
> 1.) Google now has a patent search engine out in a Beta version that let me
> find and download a PDF of it: http://www.google.com/patents
More about this in another follow-up reply...
> 2.) Some consider this to be related to the "Blackstar" TSTO vehicle AW&ST had
> the cover story about:
> http://www.aviationweek.com/aw/generic/story_generic.jsp?channel=awst&id=news/030606p1.xml
> ...being Boeing's entry in a competition to develop a system like this that
> was won by another aerospace firm (Lockheed or Rockwell).
>
> Pat
I believe this is related to 'Blackstar', or was at least a proposal laid out by
Boeing. Although I wonder if it's possible the work on the orbiter proceeded
apace ahead of this and may have been what was then thought to be 'Aurora',
with Boeing doing a bit of catch-up to retro-fit a carrier aircraft to what was
already being proposed for a small two-man orbiter. That's nothing but
sheer speculation on my part, but would explain somewhat why this patent seems
to go into much greater detail about the carrier aircraft than the orbiter,
including the complex cross-feed fueling system. (A way to be compatible with
an already or nearly complete orbiter design based on LH2/LO2/SSME?).
Other interesting ideas put forth in the patent:
The strut based release system Figs. 13-16.
On page 11 sections 1-20:
"... This aspect of the invention preferably further includes retracting the
landing gear of the transatmospheric vehicle after connecting the aircraft and
vehicle [orbiter] together and securing the vehicle in position,
fueling the vehicle by cross-feeding fuel from the aircraft to the vehicle
after retracting the landing gear of the vehicle, and operating the air
breathing engine to accomplish takeoff after fueling the vehicle. These steps
enhance the advantage of minimizing the structural requirements for the
landing gear of the vehicle. Such landing gear need only be sized for rolling
the vehicle on the ground and landing the vehicle when the vehicle is not
carrying the weight of fuel. This minimizes the structural requirements for
the landing gear and, thus, makes it possible to minimize the weight of the
landing gear. This preferred procedure also helps to further simplify launch
preparation and reduce turn around time."
Which also implies that if the orbiter is forced to make an emergency abort
with fuel still aboard, it has to be dumped during descent to insure a safe
landing. Details, details...
The final paragraph in this AVLeak article suggests the possibility that
several aerospace contractors (including *both* Boeing and Lockheed) may have
collaborated on the Blackstar TSTO system.
It also suggests that work on Blackstar may have been billed to either or both
NASP and the Navy's A-12 fighter, since both "multi-billion dollar programs
were canceled with little but technology development gains to show for massive
expenditures".
Dave
David Spain
> Two things about this patent:
> 1.) Google now has a patent search engine out in a Beta version that let me
> find and download a PDF of it: http://www.google.com/patents
>
>
>> Two things about this patent:
>> 1.) Google now has a patent search engine out in a Beta version that let
>> me find and download a PDF of it: http://www.google.com/patents
>
>worth of related vehicle patents, most of which I've never seen before.
Grist for the mill!
Dave
Pat Flannery
It would certainly save a lot of trouble if you could just roll the
orbiter under the carrier aircraft, and hoist it into position in the
belly recess.
> fueling the vehicle by cross-feeding fuel from the aircraft to the vehicle
> after retracting the landing gear of the vehicle, and operating the air
> breathing engine to accomplish takeoff after fueling the vehicle. These steps
> enhance the advantage of minimizing the structural requirements for the
> landing gear of the vehicle. Such landing gear need only be sized for rolling
> the vehicle on the ground and landing the vehicle when the vehicle is not
> carrying the weight of fuel. This minimizes the structural requirements for
> the landing gear and, thus, makes it possible to minimize the weight of the
> landing gear. This preferred procedure also helps to further simplify launch
> preparation and reduce turn around time."
>
> Which also implies that if the orbiter is forced to make an emergency abort
> with fuel still aboard, it has to be dumped during descent to insure a safe
> landing. Details, details...
>
> The final paragraph in this AVLeak article suggests the possibility that
> several aerospace contractors (including *both* Boeing and Lockheed) may have
> collaborated on the Blackstar TSTO system.
>
> It also suggests that work on Blackstar may have been billed to either or both
> NASP and the Navy's A-12 fighter, since both "multi-billion dollar programs
> were canceled with little but technology development gains to show for massive
> expenditures".
The concept of carrying a spaceplane this way goes clean back to the
B-70 program where a derivative version was to carry a spaceplane
semi-internally in a bulged belly housing.
Pat
hal...@aol.com
> On 3/25/2010 6:08 PM, David Spain wrote:
>
>
>
>
>
> >> Two things about this patent:
> >> 1.) Google now has a patent search engine out in a Beta version that let me
> >> find and download a PDF of it:http://www.google.com/patents
>
> > More about this in another follow-up reply...
>
> >> 2.) Some consider this to be related to the "Blackstar" TSTO vehicle AW&ST had
> >> the cover story about:
>
> - Show quoted text -- Hide quoted text -
>
> - Show quoted text -
theres lots of advantages for people hauling to orbit.
so nasa or contractors build a mini air launched shuttle, and a big
heavy lifter, or perhaps better a shuttle C launcher using as much of
the existing shuttle infrastructure as possible.
probably everyone would like this, a winner for existing shuttle
contractors, KSC workforce, contractors everywhere..........
heck even the budget people would likely be pleased.
existing shuttle could get reseign for unmanned launches and
landings,
to be used for large ISS components both up and down during the
transition phase..
a shuttles could go to museums eventually nasa retaing one for special
purposes:)
has anyone calculated the costs of the Killer layoffs with the
shuttles program end?
its possible keeping a reseigned program flying may be cheaper than
the layoffs nationwide, when you figure in the colateral
damages.............
unemployment, medicade, home foreclosures, all combined support
service costs for those unemployeed.
it might be enough to keep nasa operational to save money:)
now wouldnt THAT be a novel idea:)???
.......
David Spain
purports to be a cross section of a proposed scramjet engine for the
'preferred second embodiment of vehicle 50'.
As drawn there is insufficient bypass to the supersonic flow.
If all intake air *must* pass through a shockwave to the injectors
and combuster, how is this any different from a traditional ramjet?
Which limits top speeds to around Mach 3?
The cross section shown *might* have a chance to work, if what is
left out is the *huge* amount of bypass flow *around* (in front of
and behind) what is shown in the cross section.
The scramjet designs I've seen to-date that seem to have an inkling
of working only induce a shockwave in portions of the flow. Typically
shaping the shockwaves to allow combustion in the void, with the bulk
of the flow passing straight through the system unchanged. All designs
attempt to use a shaped void to pass combustion gasses through where
thrust is obtained by expansion between the engine casing and the flow
or directly expanding against the flow itself. In the first case the
engine casing and the flow act as a nozzle or nozzle ring, in the latter
case the supersonic flow itself acts as the nozzle.
Dave
Marvin the Martian
> On 3/25/2010 6:43 AM, Marvin the Martian wrote:
>>> That was the concept they were going to use on the 747-launched Air
>>> Force minishuttle of the 1980's to eliminate the need for foam
>>> insulation on its big drop tank:
>>> http://danielmarin.blogspot.com/2008/12/maks-y-alsv.html This Rockwell
>>> study would appear to carry the two-stage vehicle inside of a humidity
>>> controlled interior of a C-5B to accomplish the same prevention of ice
>>> build-up: http://www.thespacereview.com/archive/1591a.jpg
>>
>> And it was never done because it's STUPID.
>
> If it was stupid, it's a stupid that both the US and Russia put a lot of
> effort into studying over the years.
And that's how humans launch rockets to LEO now, from airplanes. Oh wait!
You don't do that. You launch them from the ground.
> One advantage you get is you can fly out over the ocean and launch into
> any orbital inclination you want without worrying about stages coming
> down in inhabited areas.
Using a barge, like Sea-Launch, would be much more practical and Sea
Launch went bankrupt. Any other stupid and failed ideas you want to warm
over and serve back up after allowing them to compost a while?