 |
|
| Science Forum Index » Physics Forum » A kerosene-fueled X-33 as a single stage to orbit... |
|
Page 3 of 4 Goto page Previous 1, 2, 3, 4 Next |
|
| Author |
Message |
| kT... |
 Posted: Mon Nov 02, 2009 11:20 am |
|
|
|
Guest
|
Jeff Findley wrote:
[quote]"Sam Wormley" <swormley1 at (no spam) mchsi.com> wrote in message
news:B_fHm.110690$5n1.67859 at (no spam) attbi_s21...
Single Stage to Orbit really limits payload "weight".
[/quote]
What you *really* want to optimize is cost per kg to orbit, not
[quote]the payload mass fraction of your vehicle.
[/quote]
Cost per kilogram doesn't give you much engineering directive.
You may as well have an engineering competition. |
|
|
| Back to top |
|
|
|
| Pat Flannery... |
| Posted: Mon Nov 02, 2009 11:25 am |
|
|
|
Guest
|
Sylvia Else wrote:
[quote]
I wonder how far aviation would have got if the first aircraft had been
required to be economically viable.
[/quote]
When designing the Flyer, it was the intention of the Wright brothers to
make a lot of money by selling them to the US and other governments for
use as military reconnaissance devices.
That's the reason they took out so many patents on it, as it was their
intention to lock any competitors out of the profits that could be
accrued that way.
They were doing pretty good at that tactic until Glenn Curtiss developed
aerilons to replace the Wright wing-warping system for roll control, and
that was not judged as a infringement of the Wright patents.
Pat |
|
|
| Back to top |
|
|
|
| Pat Flannery... |
| Posted: Mon Nov 02, 2009 11:35 am |
|
|
|
Guest
|
Robert Clark wrote:
[quote]The very key aspect of this proposal is that the tanks remain the
*same* size, but at a *lighter* weight. In fact the intent was to keep
the same shape of the X-33 and just switch out the propellant tanks
and engines. So in fact the vehicle becomes lighter for its volume
with hydrocarbon fuels.
[/quote]
The mixture rate by volume is way different for LOX/LH2 and
Kerosene/LOX, so the tanks well have to be changed in proportional size
to each other.
Also, due to the far higher density of the Kerosene versus LH2, during
acceleration a tank strong enough to carry LH2 will rupture under the
higher weight of Kerosene.
The whole point of the composite tanks on VentureStar was to get their
weight down to lower than ones made out of aluminum, so I'm keen to see
what you are going to make them out of that is lighter than composite
materials.
[quote]When you consider the other benefits of hydrocarbon fuels over
hydrogen, the higher Isp of hydrogen/LOX propellant becomes less of an
advantage.
In fact, kerosene is not necessarily the best hydrocarbon to use,
which I'll discuss in a following post.
[/quote]
I'm breathless with anticipation.
Pat |
|
|
| Back to top |
|
|
|
| Sylvia Else... |
| Posted: Mon Nov 02, 2009 8:22 pm |
|
|
|
Guest
|
Jonathan wrote:
[quote]The military
wants a missile defense base on the south pole of the Moon, which
is the only place the Earth can be continually observed.
[/quote]
Seems a long way out for a defense base that can in any case only see
just under 50% of the Earth at a time, and has most of Earth out of view
for nearly 13 consecutive hours in each 25.
Sylvia. |
|
|
| Back to top |
|
|
|
| Jonathan... |
| Posted: Tue Nov 03, 2009 12:02 am |
|
|
|
Guest
|
"Sam Wormley" <swormley1 at (no spam) mchsi.com> wrote in message
news:BZpHm.111330$5n1.96555 at (no spam) attbi_s21...
[quote]Sylvia Else wrote:
Sam Wormley wrote:
Single Stage to Orbit really limits payload "weight".
Why does that matter? The point of an SSTO is to get down the cost to orbit
per payload kg. An SSTO is likely to mass more than a disposable multi-stage
for a given payload, but that is not in itself a source of concern.
Sylvia.
http://en.wikipedia.org/wiki/Single-stage-to-orbit#SSTO_Cons
[/quote]
"The development on the RS-2200 was formally halted in early 2001
when the X-33 program did not receive Space Launch Initiative funding."
http://en.wikipedia.org/wiki/Aerospike_engine
Let's see, what happened in early 2001 to change NASA's direction
from SSTO to Lockheed's wet-dream of an endless production line
of shiny new Saturn V's on steroids???
Oh yes, President Bush took office in Jan, 2001.
From that day on, the terms...ssto, spaceports and space solar power
vanished from the NASA vocabulary. Does anyone really believe
that sea-change in goals had anything at all to do with the merits
of the technology?
President Bush did the same thing with NASA he did everywhere.
He asked the big businesses involved what their wish-list was, and
that became policy. Lockheed wanted an end to SSTO because
it would send the launch industry business elsewhere. The military
wants a missile defense base on the south pole of the Moon, which
is the only place the Earth can be continually observed.
And Voila, like magic, The Vision! Building a lunar colony on the
south pole of the Moon became policy. Which requires one-shot
(not reusable) hardware which is far /too large and sophisticated /
for the start-ups.
A marriage made in George Bush's corrupt corporate heaven.
Let's recap Lockheed stock prices around the relevant dates.
When Bush took office in Jan 01
LM stock was about $35/share.
When Bush announced the Vision in Jan 04,
LM stock was about $50/share
In January of 06, it peaked at $135/share.
400% in just five short years,
I used to wonder why VP Cheney's wife was put on the Lockheed
Board of Directors. Hell, that wasn't nearly enough payback for the
amount Lockheed benefited from Bush/Cheney policy.
And the Vision came from Cheney btw.
And in case you still doubt Bush had anything to do with Lockheed
stock prices, in November 08, just as Bush left office, LM stock
plummeted as far as $15 a share.
http://bigcharts.marketwatch.com/quickchart/quickchart.asp?symb=lm&sid=0&o_symb=lm&freq=2&time=13
What about Thiokole, another insider protected company?
Now called ATK?
Jan 01 ...$30/share
Jan 04.....$60/share
Jan 06.....$120/share
Nov 08....$75/share
Same 400% rise .....funny about that.
http://bigcharts.marketwatch.com/quickchart/quickchart.asp?symb=atk&sid=0&o_symb=atk&freq=2&time=13&x=35&y=8
Jonathan
s |
|
|
| Back to top |
|
|
|
| Jonathan... |
| Posted: Wed Nov 04, 2009 12:01 am |
|
|
|
Guest
|
"Sylvia Else" <sylvia at (no spam) not.at.this.address> wrote in message
news:006b57ff$0$26922$c3e8da3 at (no spam) news.astraweb.com...
[quote]Jonathan wrote:
The military
wants a missile defense base on the south pole of the Moon, which
is the only place the Earth can be continually observed.
Seems a long way out for a defense base that can in any case only see just
under 50% of the Earth at a time, and has most of Earth out of view for nearly
13 consecutive hours in each 25.
[/quote]
Yes, but a rocket plume in space can't be tracked from the ground.
And when the Chinese shot down that satellite, it showed anything
in low orbit can be taken out in a first strike.
Only the Moon can provide a secure ground based facility that
can track incoming missiles.
And remember, the LRO Moon mapper is a military satellite.
So the Pentagon is suddenly interested in pure exploration???
For the sake of scientific curiosity alone?
Our military is far more ambitious since 9/11 than most people
appreciate.
At the /very same time/ The Vision for Space Exploration and
it's twenty year plan was announced, the Space Command
announced it's /twenty year plan/
US Space Command
Vision for 2020
"The increasing reliance of US military forces upon space
power combined with the explosive proliferation of global
space capabilities makes a space vision essential."
http://www.fas.org/spp/military/docops/usspac/visbook.pdf
A few of our Space Wings stated 'ambitions', in their
own words.
10/28/2009 - SCHRIEVER AIR FORCE BASE, Colo. --
The 50th Space Wing's emblem contains the phrase "Master of Space".
It's difficult to make a case against that claim..."
http://www.afspc.af.mil/news/story.asp?id=123174907
460th Space Wing Mission
"Deliever global infrared surveillance, tracking and
missile warning...and provide combatant commanders
with expeditionary warrior Airmen"
http://www.buckley.af.mil/
SPACE AND MISSILE SYSTEMS CENTER
"The Space and Missile Systems Center (SMC), a subordinate unit
of the Air Force Space Command at Peterson Air Force Base,
Colorado"
"The center has an annual total budget in excess of $10 billion per year
It manages between $50 and $60 billion in contracts at
any one time."
http://www.losangeles.af.mil/library/factsheets/factsheet.asp?id=5318
Space Logistics Group
"Located at Peterson AFB, the Space Logistics Group has 550
people and a $500 million annual budget. It sustains and modifies
worldwide USAF/DoD space weapon systems..."
http://www.losangeles.af.mil/units/
SPACE INNOVATION AND DEVELOPMENT CENTER
Mission
"The Space Innovation and Development Center at Schriever Air
Force Base, Colo., is "unlocking the potential" as premier innovators
integrators and operational testers of air, space and cyberspace
power to the warfighter. The center's mission is to advance
full-spectrum warfare..."
http://www.afspc.af.mil/library/factsheets/factsheet.asp?id=3651
14th Air Force (Air Forces Strategic)
"The 14th Air Force (Air Forces Strategic) is responsible for the
organization, training, equipping, command and control, and
employment of Air Force space forces to support operational
plans and missions for U.S. combatant commanders..."
Featured Units;
21st Space Wing
30th Space Wing
14th AF IMA Link
45th Space Wing
50th Space Wing
460th Space Wing
http://www.vandenberg.af.mil/units/14thairforce.asp
Space Command boss talks of space, cyber connection
Posted 9/29/2009
66th Air Base Wing Public Affairs
9/29/2009 - LEDYARD, Conn. (AFNS) -- Addressing the
Air Force Command and Control, Intelligence, Surveillance and Reconnaissance
Symposium here Sept. 29, the leader of the
Air Force Space community said..."
"Space allows us to operate in small groups in distributed ways,"
the general said. "In near-peer conflict, space allows us to complete
the kill chain. In global assessment, it's the unblinking, or almost
unblinking, eye. In crisis management, it allows us to see what we'd
otherwise miss. ...
"It allows us to navigate with accuracy, to communicate with
certainty, strike with precision and see with clarity.
Those are enormous war fighting advantages."
http://www.24af.af.mil/news/story.asp?id=123170204
Missile Defense Agency
http://www.mda.mil/mdalink/html/mdalink.html
Directed Energy Directorate (love that name btw)
"The Air Force Research Laboratory's Directed Energy Directorate,
Kirtland AFB, N.M. is the U.S. Air Force's center of expertise
in the range of technologies required for high-energy lasers,
high-power microwaves, high-power millimeter waves
and advanced optics."
http://www.kirtland.af.mil/afrl_de/
AFRL DIRECTED ENERGY DIRECTORATE PRODUCT LINES
The Directed Energy Directorate concentrates on five major
"product" lines or emphasis areas - technologies that can fulfill
a wide range of defense needs.
Counter Electronics - Disrupt adversaries' critical military and
infrastructure electronics and communications equipment
with little to no collateral damage utilizing radio frequency
and high power microwaves
Force Protection - Protect U.S. Forces with directed energy shields
and non-lethal weaponry to minimize casualties and reduce collateral
damage
Long Range Strike - Identify, communicate and attack time critical
targets anytime; anywhere
Precision Engagement - Provide scaleable effects from disrupt to
destroy on a wide range of tactical targets with limited collateral
damage
Space Control - Provide rapid knowledge of space situations
for the combatant commander to ensure freedom of action in space
http://www.kirtland.af.mil/library/factsheets/factsheet.asp?id=7971
This is of course a partial list, and what is ....publicly known.
Pentagon's Black Budget Grows to More Than $50 Billion
"It makes the Pentagon's secret operations, including the
intelligence budgets nested inside, "roughly equal in magnitude
to the entire defense budgets of the UK, France or Japan,"
Sweetman adds. All in all, about seven and a half percent
of the Defense Department's total spending is now classified."
http://www.wired.com/dangerroom/2009/05/pentagons-black-budget-grows-to-more-than-50-billion/
The 'vision' of orbiting ships zapping the enemy from orbit
ala Star Trek isn't just some future dream. These guys
are furiously working to make it happen as soon as possible
[quote]
Sylvia.[/quote] |
|
|
| Back to top |
|
|
|
| Robert Clark... |
| Posted: Fri Nov 13, 2009 2:58 pm |
|
|
|
Guest
|
The proposal was to transform the X-33 into a reusable, orbital
vehicle using NK-33 engines. The NK-33 was a Russian 1960's era engine
so I thought they would have to be taken out of mothballs for the
purpose. But I recently found that Aerojet is working with the NK-33
engines to be used on Orbital Sciences’ Taurus 2 launcher:
08/31/09 10:15 AM ET
Aerojet Looking to Restart Production of NK-33 Engine.
By Amy Klamper
http://www.spacenews.com/launch/aerojet-looking-restart-production-nk-33-engine.html
August 19, 2009
Russian Mail-Order Ride.
http://blogs.airspacemag.com/daily-planet/2009/08/19/russian-mail-order-ride/
Aerojet has already purchased several of the engines, and is debating
whether to start it's own production lines or to use Russian
production for future purchases of the engine. Then the Air Force or
NASA could use these to make a reusable, single-stage-to-orbit vehicle
and near term, at least for a prototype vehicle. There was some debate
on the Augustine commission if NASA and the U.S. should use Russian
engines for a significant portion of their launches, but this
complaint might be ameliorated in regards to the NK-33 if production
lines were started in the U.S.
A question would be of the payload it could carry. The preliminary
calculations I made suggested it might just make orbit, so likely it
would have low payload capability. Some possibilities to increase the
payload might be to densify the kerosene propellant by subcooling to
near LOX temperatures or to use more energetic hydrocarbon
propellants. The densification would allow it carry more propellant.
Some possible energetic hydrocarbon propellants are suggested here:
Alternate Propellants for SSTO Launchers.
Dr. Bruce Dunn
http://www.dunnspace.com/alternate_ssto_propellants.htm
Another possible method to increase payload would to use a version of
the NK-33 with an aerospike nozzle. This would allow it to have higher
Isp at sea level.
It should also be possible to use the hydrocarbon fueled X-33 as the
a reusable first stage booster. The Air Force is investigating such
boosters as a means of cutting costs to space. Since the reconfigured
X-33 would be able to reach orbit at 21,700 kg dry mass, it could be
able to lift in the range of a few thousand kg's payload as the first
stage of two stage-to-orbit-system.
Bob Clark
On Nov 1, 8:20 am, Robert Clark <rgregorycl... at (no spam) yahoo.com> wrote:
[quote]Table of Contents.
I.)Introduction.
II.)Lightweight propellant tanks.
III.)Kerosene fuel and engines for the X-33/Venture star.
IVa.)Aerodynamic lift applied to ascent to orbit.
b.)Estimation of fuel saving using lift.
V.)Kerosene fueled VentureStar payload to orbit.
I.) A debate among those questing for the Holy Grail of a reusable,
single-stage-to-orbit vehicle is whether it should be powered by
hydrogen or a dense hydrocarbon such as kerosene. Most concepts for
such a vehicle centered on hydrogen, since a hydrogen/LOX combination
provides a higher Isp. However, some have argued that dense fuels
should be used since they take up less volume (equivalently more fuel
mass can be carried in the same sized tank) so they incur less air
drag and also since the largest hydrocarbon engines produce greater
thrust they can get to the desired altitude more quickly so they also
incur lower gravity drag loss.
Another key fact is that for dense fuels the ratio of propellant mass
to tank mass is higher, i.e., you need less tank mass for the same
mass of propellant. This fact is explored in this report:
Single Stage To Orbit Mass Budgets Derived From Propellant Density and
Specific Impulse.
John C. Whitehead
32nd AIAA/ASME/SAE/ASEE Joint Propulsion ConferenceLake Buena Vista,
FLJuly 1-3, 1996http://www.osti.gov/bridge/servlets/purl/379977-2LwFyZ/webviewable/37...
Whitehead notes that the propellant mass to tank mass ratio for
kerosene/LOX is typically around 100 to 1, while for liquid hydrogen/
LOX it's about 35 to 1, which would result in a significantly greater
dry mass for the hydrogen-fueled case just in tank weight alone. Based
on calculations such as these Whitehead concludes the best option for
a SSTO would be to use kerosene/LOX.
The case for the X-33/VentureStar is even worse because the unusual
shape of the tanks requires them to use more tank mass than a
comparably sized cylindrical tank. This is discussed here:
Space Access Update #91 2/7/00.
The Last Five Years: NASA Gets Handed The Ball, And Drops It.
"...part of L-M X-33's weight growth was the "multi-
lobed" propellant tanks growing considerably heavier than promised.
Neither Rockwell nor McD.-Douglas bid these; both used proven
circular-section tanks. X-33's graphite-epoxy "multi-lobed" liquid
hydrogen tanks have ended up over twice as heavy relative to the
weight of propellant carried as the Shuttle's 70's vintage aluminum
circular-section tanks - yet an X-33 tank still split open in test
last fall. Going over to aluminum will make the problem worse; X-
33's aluminum multi-lobed liquid oxygen tank is nearly four times as
heavy relative to the weight of propellant carried as Shuttle's
aluminum circular-section equivalent."http://www.space-access.org/updates/sau91.html
The X-33's twin liquid hydrogen tanks had a weight of 4,600 pounds
each, and the liquid oxygen tank a weight of 6,000 pounds, for total
of 15,200 pounds for the tanks:
Marshall Space Flight Center
Lockheed Martin Skunk Works
Sept. 28, 1999
X-33 Program in the Midst of Final Testing and Validation of Key
Components.http://www.xs4all.nl/~carlkop/x33.html
The weight of the propellant carried by the X-33 was supposed to be
210,000 lb. So the propellant to tank mass ratio for the X-33 was only
about 14 to 1(!). This would be a severe problem for the full-scale
VentureStar. Its gross lift off weight was supposed to be 2,186,000
lbs with a fuel weight of 1,929,000 lbs:
X-33 Advanced Technology Demonstrator.http://teacherlink.ed.usu.edu/tlnasa/OtherPRINT/Lithographs/X33.Advan...
So the VentureStar would have a dry mass of 257,000 lbs. Since the
same design would be used for the VentureStar tanks as those of the
X-33, the propellant mass to tank mass ratio would also be 14 to 1, so
the tank mass would be 138,000 lbs. But this means the empty tank mass
alone would be over half of the vehicle's dry weight (!)
It would have been extremely difficult for the VentureStar to have
made orbit with such a large weight penalty from the start. From all
accounts the weight problem with the tanks drove other problems such
as the need to add larger wings, increasing the weight problem
further. NASA wound up canceling the program when Lockheed couldn't
deliver the working liquid hydrogen tanks even at this excessive
weight. However, rather than canceling the program I believe the
better course would have been to open up competition for coming up
with alternative, creative solutions for reducing the weight of the
tanks. This would also have resolved some of the problems with the
vehicles weight growth.
II.) I have proposed one possibility for lightweighting the X-33 tanks
on this forum:
http://www.bautforum.com/space-exploration/86728-passenger-market-sub...
The idea would be to achieve the same lightweight tanks as
cylindrical ones by using multiple, small diameter, aluminum
cylindrical tanks. You could get the same volume by using varying
lengths and diameters of the multiple cylinders to fill up the volume
taken up by the tanks. The cylinders would not have to be especially
small. In fact they could be at centimeter to millimeter diameters, so
would be of commonly used sizes for aluminum tubes and pipes.
The weight of the tanks could be brought down to the usual 35 to 1
ratio for propellant to tank mass. Then the mass of the tanks on the
X-33 would be 210,000 lbs/35 = 6,000 lbs, saving 9,200 lbs off the
vehicle dry weight. This would allow the hydrogen-fueled X-33 to
achieve its original Mach 15 maximum velocity.
The same idea applied to the full-scale hydrogen-fueled VentureStar
would allow it to significantly increase its payload carrying
capacity. At a 35 to 1 ratio of propellant mass to tank mass, the
1,929,000 lbs propellant mass would require a mass of 1,929,000/35 > 55,000 lbs for the tanks, a saving of 83,000 lbs off the original tank
mass. This could go to extra payload, so from 45,000 lbs max payload
to 128,000 lbs max payload.
An analogous possibility might be to use a honeycombed structure for
the entire internal makeup of the tank. The X-33's carbon composite
tank was to have a honeycombed structure for the skin alone. Using a
honeycomb structure throughout the interior might result in a lighter
tank in the same way as does multiple cylinders throughout the
interior.
Still another method might be to model the tanks standing vertically
as conical but with a flat front and back, and rounded sides. Then the
problem with the front and back naturally trying to balloon out to a
circular cross section might be solved by having supporting flat
panels at regular intervals within the interior. The X-33 composite
tanks did have support arches to help prevent the tanks from
ballooning but these only went partially the way through into the
interior. You might get stronger a result by having these panels go
all the way through to the other side.
These would partition the tanks into portions. This could still work
if you had separate fuel lines, pressurizing gas lines, etc. for each
of these partitions and each got used in turn sequentially. A
preliminary calculation based on the deflection of flat plates under
pressure shows with the tank made of aluminum alloy and allowing
deflection of the flat front and back to be only of millimeters that
the support panels might add only 10% to 20% to the weight of the
tanks, while getting similar propellant mass to tank mass ratio as
cylindrical tank. See this page for an online calculator of the
deflection of flat plates:
eFunda: Plate Calculator -- Simply supported rectangular plate with
uniformly distributed loading.http://www.efunda.com/formulae/solid_mechanics/plates/calculators/SSS...
Note you might not need to have a partitioned tank, with separate
fuel lines, etc., if the panels had openings to allow the fuel to pass
through. These would look analogous to the wing spars in aircraft
wings that allow fuel to pass through. You might have the panels be in
a honeycomb form for high strength at lightweight that still allowed
the fuel to flow through the tank. Or you might have separate beams
with a spaces between them instead of solid panels that allowed the
fuel to pass through between the beams.
Another method is also related to the current design of having a
honeycombed skin for the composite hydrogen tanks. Supposed we filled
these honeycombed cells with fluid. It is known that pressurized tanks
can provide great compressive strength. This is in fact used to
provide some of the structural strength for the X-33 that would
otherwise have to be provided by heavy strengthening members. This
idea would be to apply fluid filled honycombed cells. However, what we
need for our pressurized propellant tanks is *tensile strength*.
A possible way tensile strength could be provided would be to use the
Poisson's ratio of the honeycombed cells:
Poisson's ratio.http://en.wikipedia.org/wiki/Poisson%27s_ratio
Poisson's ratio refers to the tendency of a material stretched in one
direction to shrink in length in an orthogonal direction. Most
isotropic solid materials have Poisson's ratio of about .3. However,
the usual hexagonal honeycombed structure, not being isotropic, can
have Poisson's ratios in the range of +1. This is mentioned in this
article about non-standard honeycombed structures that can even have
negative Poisson ratios:
Chiral honeycomb.http://silver.neep.wisc.edu/~lakes/PoissonChiral.html
However, note that from the formula for the volumetric change in the
Wikipedia Poisson's ratio page, a stretching of a material with a +1
Poisson's ratio implies a *decrease* in volume; actually this is true
for any case where the Poisson's ratio is greater than +.5. Then fluid
filled honeycombed cells would resist the stretching of tensile strain
by the resistance to volume compression. This would be present with
both gases and liquids. Gases are lighter. ...
read more »[/quote] |
|
|
| Back to top |
|
|
|
| Robert Clark... |
| Posted: Sat Nov 14, 2009 9:14 pm |
|
|
|
Guest
|
On Nov 13, 7:58 pm, Robert Clark <rgregorycl... at (no spam) yahoo.com> wrote:
[quote] The proposal was to transform the X-33 into a reusable, orbital
vehicle using NK-33 engines. The NK-33 was a Russian 1960's era engine
so I thought they would have to be taken out of mothballs for the
purpose. But I recently found that Aerojet is working with the NK-33
engines to be used on Orbital Sciences’ Taurus 2 launcher:
08/31/09 10:15 AM ET
Aerojet Looking to Restart Production of NK-33 Engine.
By Amy Klamperhttp://www.spacenews.com/launch/aerojet-looking-restart-production-nk...
August 19, 2009
Russian Mail-Order Ride.http://blogs.airspacemag.com/daily-planet/2009/08/19/russian-mail-ord...
Aerojet has already purchased several of the engines, and is debating
whether to start it's own production lines or to use Russian
production for future purchases of the engine. Then the Air Force or
NASA could use these to make a reusable, single-stage-to-orbit vehicle
and near term, at least for a prototype vehicle. There was some debate
on the Augustine commission if NASA and the U.S. should use Russian
engines for a significant portion of their launches, but this
complaint might be ameliorated in regards to the NK-33 if production
lines were started in the U.S.
A question would be of the payload it could carry. The preliminary
calculations I made suggested it might just make orbit, so likely it
would have low payload capability. Some possibilities to increase the
payload might be to densify the kerosene propellant by subcooling to
near LOX temperatures or to use more energetic hydrocarbon
propellants. The densification would allow it carry more propellant.
Some possible energetic hydrocarbon propellants are suggested here:
Alternate Propellants for SSTO Launchers.
Dr. Bruce Dunnhttp://www.dunnspace.com/alternate_ssto_propellants.htm
Another possible method to increase payload would to use a version of
the NK-33 with an aerospike nozzle. This would allow it to have higher
Isp at sea level.
It should also be possible to use the hydrocarbon fueled X-33 as the
a reusable first stage booster. The Air Force is investigating such
boosters as a means of cutting costs to space. Since the reconfigured
X-33 would be able to reach orbit at 21,700 kg dry mass, it could be
able to lift in the range of a few thousand kg's payload as the first
stage of two stage-to-orbit-system.
[/quote]
Aerojet claims their version of the NK-33 is "fully reusable":
Space Lift Propulsion.
http://www.aerojet.com/capabilities/spacelift.php
Anyone have any idea how many reuses they mean by that?
Bob Clark |
|
|
| Back to top |
|
|
|
| Pat Flannery... |
| Posted: Sun Nov 15, 2009 7:53 am |
|
|
|
Guest
|
|
| Back to top |
|
|
|
| Robert Clark... |
| Posted: Sat Nov 21, 2009 4:59 am |
|
|
|
Guest
|
On Nov 13, 7:58 pm, Robert Clark <rgregorycl... at (no spam) yahoo.com> wrote:
[quote] The proposal was to transform the X-33 into a reusable, orbital
vehicle using NK-33 engines. The NK-33 was a Russian 1960's era engine
so I thought they would have to be taken out of mothballs for the
purpose. But I recently found that Aerojet is working with the NK-33
engines to be used on Orbital Sciences’ Taurus 2 launcher:
08/31/09 10:15 AM ET
Aerojet Looking to Restart Production of NK-33 Engine.
By Amy Klamperhttp://www.spacenews.com/launch/aerojet-looking-restart-production-nk...
August 19, 2009
Russian Mail-Order Ride.http://blogs.airspacemag.com/daily-planet/2009/08/19/russian-mail-ord...
Aerojet has already purchased several of the engines, and is debating
whether to start it's own production lines or to use Russian
production for future purchases of the engine. Then the Air Force or
NASA could use these to make a reusable, single-stage-to-orbit vehicle
and near term, at least for a prototype vehicle. There was some debate
on the Augustine commission if NASA and the U.S. should use Russian
engines for a significant portion of their launches, but this
complaint might be ameliorated in regards to the NK-33 if production
lines were started in the U.S.
A question would be of the payload it could carry. The preliminary
calculations I made suggested it might just make orbit, so likely it
would have low payload capability. Some possibilities to increase the
payload might be to densify the kerosene propellant by subcooling to
near LOX temperatures or to use more energetic hydrocarbon
propellants. The densification would allow it carry more propellant.
Some possible energetic hydrocarbon propellants are suggested here:
Alternate Propellants for SSTO Launchers.
Dr. Bruce Dunnhttp://www.dunnspace.com/alternate_ssto_propellants.htm
Another possible method to increase payload would to use a version of
the NK-33 with an aerospike nozzle. This would allow it to have higher
Isp at sea level.
It should also be possible to use the hydrocarbon fueled X-33 as the
a reusable first stage booster. The Air Force is investigating such
boosters as a means of cutting costs to space. Since the reconfigured
X-33 would be able to reach orbit at 21,700 kg dry mass, it could be
able to lift in the range of a few thousand kg's payload as the first
stage of two stage-to-orbit-system.
[/quote]
The same reconfiguration of the Lockheed version of the X-33 to dense
fuels and engines to transform it into a full orbital vehicle would
also work for the other proposed half-scale suborbital demonstrators.
The McDonnell-Douglas version was essentially the DC-X, scaled
somewhat larger. See the linked image. I don't know how much the McD-D
version of the X-33 would have cost. However, according to this
Astronautix page a 1/2-scale version of the full orbital DC-Y had been
proposed, but not funded, which would have cost in the range $450
million, compared to the $60 million of the DC-X, in 1990's dollars:
DC-X2.
http://astronautix.com/lvs/dcx2.htm
This would have just below suborbital to suborbital performance, but
the price would be significantly less than the DC-Y full orbital
version of $5 billion:
DC-Y.
http://astronautix.com/lvs/dcy.htm
However, the point is some preliminary calculations show this 1/2-
scale DC-X2 should be able to carry enough dense hydrocarbon fuel
under such a reconfiguration to reach orbit. So you would be able to
get a reusable SSTO prototype at a significantly reduced price than
the $5 billion suggested for the full DC-Y vehicle program.
Bob Clark
Figure 5: X-33 Concept Art from McDonnell Douglas (Frassanito, J.,
McDonnell Douglas).
http://vorlon.case.edu/~jam64/images/SSTO/SSTO_Figure_5.jpg
taken from:
Single Stage to Orbit:
A Reliable Transport System or an Unattainable Dream?
http://vorlon.case.edu/~jam64/work/ssto.htm |
|
|
| Back to top |
|
|
|
| Robert Clark... |
| Posted: Mon Nov 23, 2009 10:42 pm |
|
|
|
Guest
|
On Nov 21, 9:59 am, Robert Clark <rgregorycl... at (no spam) yahoo.com> wrote:
[quote]On Nov 13, 7:58 pm, Robert Clark <rgregorycl... at (no spam) yahoo.com> wrote:
The proposal was to transform the X-33 into a reusable, orbital
vehicle using NK-33 engines. The NK-33 was a Russian 1960's era engine
so I thought they would have to be taken out of mothballs for the
purpose. But I recently found that Aerojet is working with the NK-33
engines to be used on Orbital Sciences’ Taurus 2 launcher:
08/31/09 10:15 AM ET
Aerojet Looking to Restart Production of NK-33 Engine.
By Amy Klamperhttp://www.spacenews.com/launch/aerojet-looking-restart-production-nk...
August 19, 2009
Russian Mail-Order Ride.http://blogs.airspacemag.com/daily-planet/2009/08/19/russian-mail-ord...
Aerojet has already purchased several of the engines, and is debating
whether to start it's own production lines or to use Russian
production for future purchases of the engine. Then the Air Force or
NASA could use these to make a reusable, single-stage-to-orbit vehicle
and near term, at least for a prototype vehicle. There was some debate
on the Augustine commission if NASA and the U.S. should use Russian
engines for a significant portion of their launches, but this
complaint might be ameliorated in regards to the NK-33 if production
lines were started in the U.S.
A question would be of the payload it could carry. The preliminary
calculations I made suggested it might just make orbit, so likely it
would have low payload capability. Some possibilities to increase the
payload might be to densify the kerosene propellant by subcooling to
near LOX temperatures or to use more energetic hydrocarbon
propellants. The densification would allow it carry more propellant.
Some possible energetic hydrocarbon propellants are suggested here:
Alternate Propellants for SSTO Launchers.
Dr. Bruce Dunnhttp://www.dunnspace.com/alternate_ssto_propellants.htm
Another possible method to increase payload would to use a version of
the NK-33 with an aerospike nozzle. This would allow it to have higher
Isp at sea level.
It should also be possible to use the hydrocarbon fueled X-33 as the
a reusable first stage booster. The Air Force is investigating such
boosters as a means of cutting costs to space. Since the reconfigured
X-33 would be able to reach orbit at 21,700 kg dry mass, it could be
able to lift in the range of a few thousand kg's payload as the first
stage of two stage-to-orbit-system.
The same reconfiguration of the Lockheed version of the X-33 to dense
fuels and engines to transform it into a full orbital vehicle would
also work for the other proposed half-scale suborbital demonstrators.
The McDonnell-Douglas version was essentially the DC-X, scaled
somewhat larger. See the linked image. I don't know how much the McD-D
version of the X-33 would have cost. However, according to this
Astronautix page a 1/2-scale version of the full orbital DC-Y had been
proposed, but not funded, which would have cost in the range $450
million, compared to the $60 million of the DC-X, in 1990's dollars:
DC-X2.http://astronautix.com/lvs/dcx2.htm
This would have just below suborbital to suborbital performance, but
the price would be significantly less than the DC-Y full orbital
version of $5 billion:
DC-Y.http://astronautix.com/lvs/dcy.htm
However, the point is some preliminary calculations show this 1/2-
scale DC-X2 should be able to carry enough dense hydrocarbon fuel
under such a reconfiguration to reach orbit. So you would be able to
get a reusable SSTO prototype at a significantly reduced price than
the $5 billion suggested for the full DC-Y vehicle program.
Bob Clark
Figure 5: X-33 Concept Art from McDonnell Douglas (Frassanito, J.,
McDonnell Douglas).http://vorlon.case.edu/~jam64/images/SSTO/SSTO_Figure_5.jpg
taken from:
Single Stage to Orbit:
A Reliable Transport System or an Unattainable Dream?http://vorlon.case.edu/~jam64/work/ssto.htm
[/quote]
Still the development cost of such a DC-X2 would be quite high in the
range
of $450 million (1990's dollars). So I was still thinking about how
small we
could make a scaled up, reconfigured DC-X to achieve orbit to the
extent that
one of the "New Space" companies could afford to build it. I noticed
that on
the DC-X there was a lot of empty space, at least according to the
diagrammatic
image on the Astronautix page:
DC-X.
http://www.astronautix.com/lvs/dcx.htm
I estimated that if we actually fully used up the conical internal
space with
propellant, with just a small area at the top for payload or no
internal
payload bay at all, made it of an all composite construction (remember
the
DC-X was not weight optimized since it would not even go suborbital)
and if
we used highly densified hydrocarbon/LOX propellant, to near the solid
phase,
then we could get quite high velocities from the DC-X, perhaps up to
Mach 20.
In that case only a small scale up from the original DC-X dimensions
would
allow you to reach full orbital performance. This would be much
cheaper than
the DC-X2. I'm thinking it might even doable for less than $100
million in
current dollars.
Then this could be doable by one of the New Space companies,
particularly
those with deep pockets such as SpaceX, Scaled Composites, XCor, Blue
Origin, etc.
The case of Blue Origin is particularly interesting because several of
the
DC-X engineers moved over to work for Blue Origin and the design of
its "New
Shepard" suborbital craft has been likened to that of the DC-X. Blue
Origin's
head Jeff Bezos has also said his intention is to move to orbital
craft:
Blue Origin.
http://en.wikipedia.org/wiki/Blue_Origin
Blue Origin New Shepard.
http://en.wikipedia.org/wiki/Blue_Origin_New_Shepard
Bob Clark |
|
|
| Back to top |
|
|
|
| Robert Clark... |
| Posted: Fri Nov 27, 2009 5:18 am |
|
|
|
Guest
|
On Nov 24, 3:42 am, Robert Clark <rgregorycl... at (no spam) yahoo.com> wrote:
[quote]...
The same reconfiguration of the Lockheed version of the X-33 to dense
fuels and engines to transform it into a full orbital vehicle would
also work for the other proposed half-scale suborbital demonstrators.
The McDonnell-Douglas version was essentially the DC-X, scaled
somewhat larger. See the linked image. I don't know how much the McD-D
version of the X-33 would have cost. However, according to this
Astronautix page a 1/2-scale version of the full orbital DC-Y had been
proposed, but not funded, which would have cost in the range $450
million, compared to the $60 million of the DC-X, in 1990's dollars:
DC-X2.http://astronautix.com/lvs/dcx2.htm
This would have just below suborbital to suborbital performance, but
the price would be significantly less than the DC-Y full orbital
version of $5 billion:
DC-Y.http://astronautix.com/lvs/dcy.htm
However, the point is some preliminary calculations show this 1/2-
scale DC-X2 should be able to carry enough dense hydrocarbon fuel
under such a reconfiguration to reach orbit. So you would be able to
get a reusable SSTO prototype at a significantly reduced price than
the $5 billion suggested for the full DC-Y vehicle program.
Bob Clark
Figure 5: X-33 Concept Art from McDonnell Douglas (Frassanito, J.,
McDonnell Douglas).http://vorlon.case.edu/~jam64/images/SSTO/SSTO_Figure_5.jpg
taken from:
Single Stage to Orbit:
A Reliable Transport System or an Unattainable Dream?http://vorlon.case..edu/~jam64/work/ssto.htm
Still the development cost of such a DC-X2 would be quite high in the range
of $450 million (1990's dollars). So I was still thinking about how small we
could make a scaled up, reconfigured DC-X to achieve orbit to the extent that
one of the "New Space" companies could afford to build it. I noticed that on
the DC-X there was a lot of empty space, at least according to the diagrammatic image on the Astronautix page:
DC-X.http://www.astronautix.com/lvs/dcx.htm
I estimated that if we actually fully used up the conical internal
space with propellant, with just a small area at the top for payload or no internal
payload bay at all, made it of an all composite construction (remember the
DC-X was not weight optimized since it would not even go suborbital) and if
we used highly densified hydrocarbon/LOX propellant, to near the solid phase,
then we could get quite high velocities from the DC-X, perhaps up to Mach 20.
In that case only a small scale up from the original DC-X dimensions would
allow you to reach full orbital performance. This would be much cheaper than
the DC-X2. I'm thinking it might even doable for less than $100 million in
current dollars. Then this could be doable by one of the New Space companies, particularly those with deep pockets such as SpaceX, Scaled Composites, XCor, Blue Origin, etc.
The case of Blue Origin is particularly interesting because several of the
DC-X engineers moved over to work for Blue Origin and the design of its "New
Shepard" suborbital craft has been likened to that of the DC-X. Blue Origin's
head Jeff Bezos has also said his intention is to move to orbital craft:
Blue Origin.http://en.wikipedia.org/wiki/Blue_Origin
Blue Origin New Shepard.http://en.wikipedia.org/wiki/Blue_Origin_New_Shepard
[/quote]
One of the well-financed New Space companies could develop a small-
payload capable, all composite, dense propellant VTVL SSTO. This might
still cost ca. $100 million. So perhaps just like NASA wanted a half-
scale suborbital demonstrator first, perhaps the New Space companies
could do this as well.
That is, since a slightly larger all-composite, weight optimized,
dense propellant DC-X might be orbit-capable, perhaps the New Space
companies could do a half-scale version of *this*. This should be
capable of high Mach, hypersonic velocities and suborbital flight. We
might estimate that since the size would be 1/2-scale, the volume and
mass might be 1/8, and the cost might therefore be 1/8 of the dense
propellant version of the DC-X so in the range of $12 million. This
might be an amount the New Space companies might want to take a chance
on.
But it would be really great if even the small New Space companies
could also investigate this. I'm thinking of companies for example
like Armadillo Aerospace and Masten Space Systems that took part in
the Lunar Lander X-prize competition. I've read that the costs of
carbon fiber composites are dropping markedly, so much so that soon
some passenger cars will be brought to market with carbon composites
making up a significant portion of their mass, something that
previously was restricted to million dollar race cars.
So some of these smaller companies might be able to make some small
test vehicles using all composite construction that would confirm the
principle that all composite construction can result in such large
mass ratios that it is equivalent to having SSTO performance. For such
small test vehicles these would not need to be reusable, so could save
weight on landing gear, thermal protection, wings or stored propellant
for landing, etc. These would just be proof of principle concept
vehicles that would suggest that with proper scaling relationships a
larger all composite vehicle should be SSTO and reusable. See the
discussion of the scaling relationships of orbital vehicles here:
Reusable launch system.
2 Reusability concepts.
2.1 Single stage.
http://en.wikipedia.org/wiki/Reusable_launch_system#Single_stage
These New Space companies might be able to keep the costs for these
small-scale demonstrators low by doing something I hadn't previously
known was possible: making your own carbon composite structures in
house.
After a web search I saw that some amateurs use carbon composites to
save weight both for home-built aircraft and model aircraft:
Homebuilt aircraft.
http://en.wikipedia.org/wiki/Homebuilt_aircraft#Composite
Carbon Fiber Composites.
http://winshiprc.tripod.com/carbon_fiber_composites.htm
Then by making their composite structures in house the New Space
companies could reduce their costs significantly at least for these
small scale test vehicles.
Bob Clark |
|
|
| Back to top |
|
|
|
| Robert Clark... |
| Posted: Fri Nov 27, 2009 6:15 am |
|
|
|
Guest
|
On Nov 21, 9:59 am, Robert Clark <rgregorycl... at (no spam) yahoo.com> wrote:
[quote]...
The same reconfiguration of the Lockheed version of the X-33 to dense
fuels and engines to transform it into a full orbital vehicle would
also work for the other proposed half-scale suborbital demonstrators.
The McDonnell-Douglas version was essentially the DC-X, scaled
somewhat larger. See the linked image. I don't know how much the McD-D
version of the X-33 would have cost. However, according to this
Astronautix page a 1/2-scale version of the full orbital DC-Y had been
proposed, but not funded, which would have cost in the range $450
million, compared to the $60 million of the DC-X, in 1990's dollars:
DC-X2.http://astronautix.com/lvs/dcx2.htm
This would have just below suborbital to suborbital performance, but
the price would be significantly less than the DC-Y full orbital
version of $5 billion:
DC-Y.http://astronautix.com/lvs/dcy.htm
However, the point is some preliminary calculations show this 1/2-
scale DC-X2 should be able to carry enough dense hydrocarbon fuel
under such a reconfiguration to reach orbit. So you would be able to
get a reusable SSTO prototype at a significantly reduced price than
the $5 billion suggested for the full DC-Y vehicle program.
Bob Clark
Figure 5: X-33 Concept Art from McDonnell Douglas (Frassanito, J.,
McDonnell Douglas).http://vorlon.case.edu/~jam64/images/SSTO/SSTO_Figure_5.jpg
taken from:
Single Stage to Orbit:
A Reliable Transport System or an Unattainable Dream?http://vorlon.case.edu/~jam64/work/ssto.htm
[/quote]
Guys, it's a simple equation to see why a reusable SSTO vehicle
should be possible.
It has been often noted that the 1960's era Titan II first stage in
itself had single-stage-to-orbit performance, though it would have had
minimal payload capability:
SSTO Cons.
"A SSTO vehicle needs to lift its entire structure into orbit. To
reach orbit with a useful payload, the rocket requires careful and
extensive engineering to save weight. This is much harder to design
and engineer. A staged rocket greatly reduces the total mass that
flies all the way into space; the rocket is continually shedding fuel
tanks and engines that are now dead weight.
"Although a SSTO rocket might theoretically be built, margins would be
likely to be very thin- even comparatively minor problems may tend to
mean that a project to achieve this could fail to achieve the
necessary mass-fraction to reach orbit with useful payload.
"Single-stage rockets were once thought to be beyond reach, but
advances in materials technology and construction techniques have
shown them to be possible. For example, calculations show that the
Titan II first stage, launched on its own, would have a 25-to-1 ratio
of fuel to vehicle hardware.[1] It has a sufficiently efficient engine
to achieve orbit, but without carrying much payload.[2]"
http://en.wikipedia.org/wiki/Single-stage-to-orbit#SSTO_Cons
See the section on Titan II first stage fully fueled mass and empty
mass here:
Titan.
"Stage1: 1 x Titan 2-1. Gross Mass: 117,866 kg (259,850 lb). Empty
Mass: 6,736 kg (14,850 lb). Motor: 2 x LR87-7. Thrust (vac): 2,172.231
kN (488,337 lbf). Isp: 296 sec. Burn time: 139 sec. Length: 22.28 m
(73.09 ft). Diameter: 3.05 m (10.00 ft). Propellants: N2O4/
Aerozine-50."
http://www.astronautix.com/lvs/titan.htm
The two LR-87-7 engines used had a mass of 713 kg each, for a total of
1426 kg:
LR87.
"Engine Model: LR87-7. Manufacturer Name: AJ23-134. Government
Designation: LR87-7. Designer: Aerojet. Propellants: N2O4/Aerozine-50.
Thrust(vac): 1,086.100 kN (244,165 lbf). Thrust(sl): 946.700 kN
(212,827 lbf). Isp: 296 sec. Isp (sea level): 258 sec. Burn time: 139
sec. Mass Engine: 713 kg (1,571 lb). Diameter: 1.53 m (5.00 ft).
Length: 3.13 m (10.26 ft). Chambers: 1. Chamber Pressure: 47.00 bar.
Area Ratio: 9.00. Oxidizer to Fuel Ratio: 1.90. Thrust to Weight
Ratio: 155.33. Country: USA. Status: Study 1961. First Flight: 1962.
Last Flight: 2003. Flown: 212."
http://www.astronautix.com/engines/lr87.htm
Most of the remaining empty mass of 5310 kg for the Titan II first
stage would be structural mass, the propellant tanks, support
structures, etc. This would be primarily aluminum and steel. Since a
1/3 to 1/2 weight saving can be made over aluminum and steel by using
carbon composites, an all composite construction could save at least
1,770 kg off the vehicle empty mass.
However, probably we would have to swap out the engine because as
described here, the LR-87-7 engine was not throttlable, which would be
needed for a SSTO:
--------------------------------------------------------------
Newsgroups: sci.space.tech
From: henry at (no spam) spsystems.net (Henry Spencer)
Subject: Re: Is Roton Dead?
Date: Tue, 9 Jan 2001 21:11:51 GMT
In article <93eaqs$6a4$1 at (no spam) mulga.cs.mu.OZ.AU>,
David Kinny <dnk at (no spam) OMIT.cs.mu.oz.au> wrote:
[quote]...in fact, the central problem with using
the Titan II first stage as an SSTO is that it has *too much* thrust to
fly an efficient trajectory.
How exactly does too much thrust prevent flying an efficient trajectory?
Difficulties in flipping over to horizontal? Or something else?
[/quote]
Basically, in the time it takes to climb clear of the atmosphere, it
picks
up too much vertical velocity. This thing was an ICBM, designed to
move
out fast... and flying as an SSTO, it hasn't got a hulking great
second
stage on top to slow it down. (In fact, a secondary problem of having
too
much thrust is the bone-crushing acceleration toward the end, when the
tanks are almost empty.) An SSTO launcher wants to take things a bit
slower, so that it can tip over to horizontal gradually, as it leaves
the
atmosphere, and still have most of its fuel left for horizontal
acceleration.
You can't just throttle back the engine, first because it wasn't
throttlable  , and second because you need to keep it operating
efficiently, which throttling usually sacrifices to at least some
extent.
However, *reducing* the performance of an engine is usually not a
difficult engineering problem!
--
When failure is not an option, success | Henry Spencer henry at (no spam) ****.net
can get expensive. -- Peter Stibrany | (aka **** at (no spam) zoo.toronto.edu)
http://yarchive.net/space/launchers/ssto.html
--------------------------------------------------------------
I suggest the NK-33 be used. It was designed for kerosene/LOX but
quite likely would also work with the N2O4/Aerozine-50 propellant of
the LR-87-7 because the LR-87 engine was variously used with N2O4/
Aerozine-50 and kerosene/LOX. Using a single NK-33 would also save 200
kg off the vehicle dry weight:
NK-33.
http://www.astronautix.com/engines/nk33.htm
The question is could we use that approx. 2,000 kg saved weight for
landing gear and thermal protection to make the vehicle reusable?
Let's take the landed weight as still 6,736 kg where we used the saved
weight for landing gear, thermal protection, and payload.
The landing gear for an aerial vehicle is commonly taken as 3% of the
landed weight:
Landing gear weight.
http://yarchive.net/space/launchers/landing_gear_weight.html
So this is 202 kg.
To make a powered vertical landing the common estimate is 10% of the
vehicle landed weight has to be used in propellant:
Reusable launch system.
Vertical landing.
http://en.wikipedia.org/wiki/Reusable_launch_system#Vertical_landing
So 673 kg.
For thermal protection, we'll assume it'll make a ballistic reentry,
base first. For this vehicle the base will only be 3 meters wide, for
an area of 7 m^2. Using base first reentry we'll have to cover
primarily the base only:
Blue Origin New Shepard.
"A passenger and cargo spacecraft has considerably less need for cross-
range."
....
"As a result, the craft is much "rounder" than the DC-X, optimized for
tankage and structural benefits rather than re-entry aerodynamics. It
has not been stated if the vehicle is intended to re-enter base-first
or nose first, but the former is most likely for a variety of reasons.
For one, it reduces heat shield area, and thus weight, covering only
the smaller bottom surface rather than the much larger upper portions.
The area around the engines would likely require some sort of heat
protection anyway, so by using the base as the heat shield the two can
be combined. This re-entry attitude also has the advantage of allowing
the spacecraft to descend all the way from orbit to touchdown in a
base-first orientation, which would seem to offer some safety benefits
as well as reducing aero-loading issues."
http://en.wikipedia.org/wiki/Blue_Origin_New_Shepard
We'll use the high temperature resistant but low maintenance metallic
shingles developed for the X-33:
REUSABLE METALLIC THERMAL PROTECTION SYSTEMS DEVELOPMENT,
http://reference.kfupm.edu.sa/content/r/e/reusable_metallic_thermal_protection_sys_117853.pdf
The areal density of this is in the range of 10 to 15 kg/m^2. This
will then require 70 to 105 kg to cover the base only.
Then the total mass for landing and thermal protection is 980 kg, and
about 1,000 kg could go to payload. This would be only 0.8% of the
gross mass but would be a reusable SSTO vehicle.
It might be possible to improve this payload fraction by using
kerosene/LOX instead of the N2O4/Aerozine-50 propellant. This would
result in a higher Isp, however the N2O4/Aerozine-50 is denser and so
more fuel can be carried.
Bob Clark |
|
|
| Back to top |
|
|
|
| Robert Clark... |
| Posted: Sat Nov 28, 2009 6:24 am |
|
|
|
Guest
|
It is important to remember that single-stage-to-orbit in itself is
not impossible. It was in fact proven to be feasible from the early
days of the space program:
Single-stage-to-orbit.
SSTO Cons.
"Single-stage rockets were once thought to be beyond reach, but
advances in materials technology and construction techniques have
shown them to be possible. For example, calculations show that the
Titan II first stage, launched on its own, would have a 25-to-1 ratio
of fuel to vehicle hardware.[1] It has a sufficiently efficient engine
to achieve orbit, but without carrying much payload.[2]"
http://en.wikipedia.org/wiki/Single-stage-to-orbit#SSTO_Cons
Such a vehicle of course while carrying minimal payload would also not
be reusable. The question is could you replicate this performance
using lightweight materials so this weight savings could go to reentry
and return systems and could this be done economically?
I already gave the argument that the weight savings possible from
composite construction makes such a reusable SSTO possible. The reason
why I say it is now economically feasible is because lightweight
carbon composite construction is now being planned for some passenger
cars. Consider the price for carbon composites in the early 90's:
G.M. to Show a High-Mileage Experimental Car
By DORON P. LEVIN,
Published: Monday, December 30, 1991
"At the North American International Auto Show in Detroit next week,
G.M. will show its Ultralite, which the company says can produce 100-
mile-a-gallon fuel efficiency at 50-mile-an-hour highway speeds.
"That efficiency is possible, G.M. said, because the car weighs only
1,400 pounds. (A Chevrolet Corsica, which is approximately the same
size as the Ultralite, weighs about twice as much.) Scaled Composites
Inc. of Mojave, Calif., built the Ultralite body for G.M.
"Although many race cars are made of carbon fiber, which is quite
sturdy, the material is enormously expensive compared with steel or
aluminum. But G.M. said it had received a patent for a process that
sharply reduces the cost of carbon fiber, which currently is about $40
a pound, compared with about 35 cents a pound for steel."
http://www.nytimes.com/1991/12/30/business/gm-to-show-a-high-mileage-experimental-car.html
So the price then was about 100 times greater than steel. You wouldn't
see many all-composite-construction rockets at those prices even if
even then it would have made a reusable SSTO possible.
Now look at the price given in this article from the year 2000:
Carbon-Fiber Composites for Cars.
"To meet the ultimate PNGV mileage goal, one potentially enabling
technology is to use carbon-fiber composites, which form the structure
of U.S. fighter jets. Carbon-fiber composites weigh about one-fifth as
much as steel, but can be comparable or better in terms of stiffness
and strength, depending on fiber grade and orientation. These
composites do not rust or corrode like steel or aluminum. Perhaps most
important, they could reduce vehicle weight by as much as 60%,
significantly increasing vehicle fuel economy.
"The problem is that carbon-fiber composites cost at least 20 times as
much as steel, and the automobile industry is not interested in using
them until the price of carbon fiber drops from $8 to $5 (and
preferably $3) a pound. Production of carbon fibers is too expensive
and slow."
http://www.ornl.gov/info/ornlreview/v33_3_00/carbon.htm
Now this British company claims their patented process allows
composite construction both for the chassis frame and the body panels
at low cost for a passenger car to be introduced next year:
Axon announces affordable, 100mpg, carbon-composite passenger car.
"Axon has gone simply for an uncomplicated 500cc engine in a low-
weight body, which replaces the traditional heavy steel or aluminium
frame with recycled carbon fibre composites - as strong as steel but
only around 40% as heavy. Extensive use of carbon materials through
Axon’s cars makes a massive impact on the power-to-weight ratio,
meaning they can get acceptable overall performance using a much
smaller, lighter and more frugal engine.
"The lightness and strength of carbon fibre have been well-known for
decades - it’s been cost that’s prevented this wonder-material from
popping up all over the automotive world, restricting it to top-end
specials and aftermarket goodies. But it’s here that Axon claim to
have made a breakthrough."
http://www.transport20.com/uncategorized/axon-announces-affordable-100mpg-carbon-composite-passenger-car/
Because of the rate at which the costs of carbon composite production
is decreasing, I argue the production cost for a reusable SSTO using
carbon composite construction, because the lighter weight in materials
required, will soon be comparable to that of an expendable rocket
using standard, heavy construction materials. And it is already now
economically feasible due to lower per use costs of a reusable
vehicle.
It is also extremely important to keep in mind that such a reduction
in structural mass for a rocket would result in a comparable reduction
in engine mass. This is important because the engine mass is the
second greatest component for the dry mass of the rocket after the
structural mass.
The reason this engine mass reduction occurs is exactly analogous to
why it occurs when replacing the structural mass of cars with lighter
materials:
Carbon Fibre Reinforced Composite Car.
Primary author: Andrew Mills
Source: Materials World, Vol 10, no. 9 pp. 20-22, September 2002.
"In the area of vehicle design, body weight is the most important
target for improvement, as a reduction in the weight of a vehicle’s
body means that a smaller engine, and a lighter drive train and
assembly can be used. This ‘benign spiral’ leads to further mass
reductions, so much so that various studies have indicated a potential
for savings of up to 65% by using carbon fibre composites instead of
steel wherever possible."
http://www.azom.com/Details.asp?ArticleID=1662#_Background
Bob Clark
Axontex chasis.
http://www.transport20.com/gallery/albums/Axon/axontex_chassis.jpg |
|
|
| Back to top |
|
|
|
| Robert Clark... |
| Posted: Tue Dec 01, 2009 4:26 am |
|
|
|
Guest
|
On Nov 29, 6:48 am, Robert Clark <rgregorycl... at (no spam) yahoo.com> wrote:
[quote] LORAL Space Systems, the leading communications satellite builder,
had a design for a single-stage-to-orbit though expendable launcher.
They expected to use all-composite cryogenic tanks on these launchers
to save weight. Their idea was that the high cost of launch is from
trying to assure high reliability. However, their launchers were to be
designed to be used for payloads such as replacing consumables on the
ISS, launching propellants to orbital depots, etc.
They were able to conclude based on study of prior launchers that
high reliable launchers cost more and correspondingly lower reliable
ones cost less. They therefore specifically aimed for a rather low
reliability rate of about 66% to get low cost. They figured this would
be allowable for low cost items such fuel and consumables.
Still, it is interesting that their low cost design was specifically
based on a SSTO, composite-tank rocket:
Aquarius: Low-Cost Low Reliability Consumables Launcher.
Enabling Technology includes large, lightweight liner-less composite
tanks.http://homepage.mac.com/fcrossman/NorCalSAMPE/Comp_WS_papers/Turner_0...
Aquarius.
"Proposed expendable, water launch, single-stage-to-orbit, liquid
oxygen/hydrogen, low-cost launch vehicle designed to carry small bulk
payloads to low earth orbit. A unique attribute was that low
reliability was accepted in order to achieve low cost."http://www.astronautix.com/lvs/aquarius.htm
Bob Clark
[/quote]
Nice list of launch vehicle designs, including some SSTO's going back
to the 60's:
Space Future - Vehicle Designs.
http://www.spacefuture.com/vehicles/designs.shtml
Here's a review of SSTO concepts proposed over the years:
History of the Phoenix VTOL SSTO and Recent Developments in Single-
Stage Launch Systems.
Gary C Hudson
http://www.spacefuture.com/archive/history_of_the_phoenix_vtol_ssto_and_recent_developments_in_single_stage_launch_systems.shtml
And this article argues that SSTO performance has long been possible
for expendables, and that a reusable one is possible with modern
materials:
Launch Vehicle Design.
"Contrary to what many people who make expendable rockets will tell
you, it isn't difficult to design a "single stage to orbit" ( SSTO)
rocket. In fact it's very easy - it can be done with rocket engines
and propellant tanks designed, manufactured and operated 20 years ago!
It's important to know this, because a lot of people will try to tell
you otherwise.
"A Thought Experiment
"This very idea was written up by Gary Hudson in "A Single-Stage-to-
Orbit thought experiment".
"If you attach 6 SSMEs (Space Shuttle Main Engines) directly to a
Space Shuttle External Tank ( ET), you could launch 30 tons payload to
orbit. It wouldn't be an economical way to launch - but it's certainly
possible. But please note: it's only possible taking off vertically;
no-one can build a horizontal take-off SSTO.
"But, of course, if you carry passengers to orbit you'll want to bring
them back - and that's what's tricky: to build a fully reusable SSTO,
not an expendable, one-way ride. "
http://www.spacefuture.com/vehicles/building.shtml
As this article notes, many people don't think a SSTO vehicle is
possible even with expendables. That is why with the rapid drop in the
cost of composite materials I'm arguing that small test vehicles of
all-composite construction should be built to prove the principle of
SSTO at least for expendables. This would be possible and affordable
to do even for the smallest of the New Space companies with in house
construction of the composite materials.
Then when it is seen that SSTO, though not reusable, performance is
possible for an actual working rocket, it will be more believable that
following well-known scaling principles that larger rockets should
allow reusable versions with significant payloads.
Bob Clark |
|
|
| Back to top |
|
|
|
|
|
All times are GMT - 5 Hours
The time now is Fri Dec 11, 2009 11:42 pm
|
|