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The X-33 as the upper stage of a two-stage-to-orbit system.



 
 
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  #1  
Old September 8th 09, 02:00 PM posted to sci.astro,sci.space.policy,sci.physics,sci.space.history
Robert Clark
external usenet poster
 
Posts: 1,150
Default The X-33 as the upper stage of a two-stage-to-orbit system.

I.) In 2006, Aviation Week claimed in an article that the Air Force
had been
testing a two-stage-to-orbit fully reusable launch system:

Two-Stage-to-Orbit ''Blackstar'' System Shelved at Groom Lake?
Mar 5, 2006
By William B. Scott
http://www.aviationweek.com/aw/gener...s/030606p1.xml

TSTO spaceplanes.
http://robotpig.net/aerospace/en_tsto.php

The system, code named "Blackstar", was supposed to consist of a Mach
3
jet-powered carrier vehicle, and a two-man orbital spaceplane.
The article was skeptically received:

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/

Blackstar A False Messiah From Groom Lake.
by Jeffrey F. Bell
Honolulu HI (SPX) Mar 10, 2006
http://www.spacedaily.com/reports/Bl...room_Lake.html

But this concept is in fact currently workable using the X-33 as the
spaceplane with a modification I suggested before to lighten the fuel
tanks of using multiple cylindrical tanks.

The X-33 was supposed to have a weight of 273,000 lbs fully fueled:

X-33 VentureStar.
Overview
http://www.fas.org/programs/ssp/man/...lanes/x33.html

The carrier craft for "Blackstar" was likened to the XB-70 Valkyrie,
a 1960's prototype of a Mach 3 bomber:

XB-70 Valkyrie.
http://en.wikipedia.org/wiki/XB-70_Valkyrie

The XB-70 had a max take off weight of 550,000 lbs and an empty
weight of
210,000, leaving 340,000 lbs for fuel and payload. The max designed
payload
for the XB-70 was supposed to be only in the range of 20,000 lbs,
however.
Using this as a model for the carrier craft would require stengthening
of the
frame and removal of some of the fuel tanks so that it could only make
short
trips to high altitude and not long distance missions. This would be
analogous
to what was done with the Shuttle Carrier Aircraft, the modified 747
used to
ferry the shuttle orbiters.
SpaceShipOne showed that lightweight carbon composites could
withstand the
heating at Mach 3 speeds. The XB-70 was composed primarily of steel.
Replacing
the steel in its airframe with carbon composites could conceivably cut
its
weight to half to one-third of its original dry weight, allowing
higher
altitude or carrying capacity.
The X-33 though was meant to be only a suborbital test vehicle and
not to
carry payload. However, the image attached below from the "X-33
VentureStar"
page shows it does have a payload bay. This would in fact be large
enough to
carry a two-man cockpit, a la an F-14-style inline seating
arrangement.
How does the suborbital X-33 attain sufficient delta-v to attain
orbit with
airlaunch? Let's say the Mach 3 carrier craft at X-33 release can give
it an
initial velocity of 1,000 m/s at an altitude of 100,000 ft, 30 km, and
we
wanted to get to orbit at 100 km.
Orbital velocity is about 7,800 m/s. However, we also need to expend
fuel
to get to the desired altitude and to overcome air drag. This page
shows the delta-v
losses due to air drag are surprisingly low:

Drag: Loss in Ascent, Gain in Descent, and What It Means for
Scalability.
Ariane A-44L: Gravity Loss: 1576 m/s Drag Loss: 135 m/s
Atlas I: Gravity Loss: 1395 m/s Drag Loss: 110 m/s
Delta 7925: Gravity Loss: 1150 m/s Drag Loss: 136 m/s
Shuttle: Gravity Loss: 1222 m/s Drag Loss: 107 m/s
Saturn V: Gravity Loss: 1534 m/s Drag Loss: 40 m/s (!!)
Titan IV/Centaur: Gravity Loss: 1442 m/s Drag Loss: 156 m/s
http://gravityloss.wordpress.com/200...r-scalability/

I'll take the air drag loss in the range of 100 m/s. The gravity loss
depends
on many factors including the altitude of the orbit you want to reach.
A
common way to estimate it is to use the equation you get from setting
the
potential energy of a certain altitude equal to the kinetic energy of
the
velocity needed to attain that altitude by ballistic launch: v = sqrt
(2gh) .
So at h= 100,000 m, v = sqrt(2*9.8*100,000) = 1,400 m/s. Then the
total
delta-v required would be 7,800 + 1,400 + 100 = 9,300 m/s.
That 1,400 m/s delta-v just for the altitude of orbit, aside from the
orbital
speed required, is quite large. But key is that this can be reduced if
your
vehicle is given a high initial speed. The idea is that the equation
between
kinetic and gravitational potential energy actually looks like this:
(1/2)Vf^2 -(1/2)vi^2 = 2gh. Then if we have an initial speed of 1,000
m/s and
we only need an additional delta-v to go from our launch altitude of
30 km to
the orbital altitude of 100 km, then the equation becomes:

(1/2)(1540.1)^2 - (1/2)(1000)^2 = 2*9.8*70,000. So this means we need
and
additional delta-v of only 540 m/s to add to our initial 1,000 m/s
speed to
reach the 100 km altitude.
Then our required total delta-v is 7,800 + 540 + 100 = 8,440. Now to
see how
much delta-v needs to be supplied by the X-33 we'll also include the
fact that
by launching near the equator we get an additional 462 m/s speed for
free, so:
delta-v still needed = 8,440 - 1,000 - 462 = 6,978 m/s.

Since the altitude for our launch is so high we can get the full
vacuum Isp
of our engines. Then the mass ratio for a X-33 engine Isp of 455 s and
required delta-v of 6,978 m/s is:

exp(6978/4550) = 4.635, which means the fuel ratio is 3.635.
Here is where the lightened fuel tanks are needed. The dry mass of
the X-33
was supposed to be 65,000 lb. Then the fuel mass would be 3.635 *
65,000 =
236,275 lb. But the tanks could only hold 210,000 lb of fuel.
I discussed the proposal for lightening the X-33 tanks he

Passenger market for suborbital, hypersonic transports.
http://www.bautforum.com/space-explo...ml#post1495726

The idea is that the problem with the X-33 tanks was their unusual,
noncylindrical shape, which means more tank mass has to be used to
hold the fuel than for usual cylindrical tanks. The suggestion was to
use multiple cylindrical tanks to get the same storage of fuel at a
low weight. This would cut the mass of the two liquid hydrogen tanks
from 9,200 lb. total to 4,600 lb. total, and would cut the mass of the
liquid oxygen tank from 6,000 lb to 1,500 lb, for a total weight
saving of 9,100 lb. Then the dry weight would become 55,900 lb., and
the fuel weight, 203,200 lb, which is within the fuel capacity of the
X-33.



Bob Clark

Image of interior of the X-33.
http://www.fas.org/spp/guide/usa/lau...3_interior.jpg
  #2  
Old September 11th 09, 05:18 PM posted to sci.astro,sci.space.policy,sci.physics,sci.space.history
Robert Clark
external usenet poster
 
Posts: 1,150
Default The X-33 as the upper stage of a two-stage-to-orbit system.

On Sep 8, 9:00*am, Robert Clark wrote:
I.) In 2006, Aviation Week claimed in an article that the Air Force
had been
testing a two-stage-to-orbit fully reusable launch system:

Two-Stage-to-Orbit ''Blackstar'' System Shelved at Groom Lake?
Mar 5, 2006
By William B. Scotthttp://www.aviationweek.com/aw/generic/story_generic.jsp?channel=awst...

TSTO spaceplanes.http://robotpig.net/aerospace/en_tsto.php

*The system, code named "Blackstar", was supposed to consist of a Mach
3
jet-powered carrier vehicle, and a two-man orbital spaceplane.
*The article was skeptically received:

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, 2006http://www.msnbc.msn.com/id/11691989/

Blackstar A False Messiah From Groom Lake.
by Jeffrey F. Bell
Honolulu HI (SPX) Mar 10, 2006http://www.spacedaily.com/reports/Blackstar_A_False_Messiah_From_Groo...

*But this concept is in fact currently workable using the X-33 as the
spaceplane with a modification I suggested before to lighten the fuel
tanks of using multiple cylindrical tanks.

*The X-33 was supposed to have a weight of 273,000 lbs fully fueled:

X-33 VentureStar.
Overviewhttp://www.fas.org/programs/ssp/man/uswpns/air/xplanes/x33.html

*The carrier craft for "Blackstar" was likened to the XB-70 Valkyrie,
a 1960's prototype of a Mach 3 bomber:

XB-70 Valkyrie.http://en.wikipedia.org/wiki/XB-70_Valkyrie

*The XB-70 had a max take off weight of 550,000 lbs and an empty
weight of
210,000, leaving 340,000 lbs for fuel and payload. The max designed
payload
for the XB-70 was supposed to be only in the range of 20,000 lbs,
however.
Using this as a model for the carrier craft would require stengthening
of the
frame and removal of some of the fuel tanks so that it could only make
short
trips to high altitude and not long distance missions. This would be
analogous
to what was done with the Shuttle Carrier Aircraft, the modified 747
used to
ferry the shuttle orbiters.
*SpaceShipOne showed that lightweight carbon composites could
withstand the
heating at Mach 3 speeds. The XB-70 was composed primarily of steel.
Replacing
the steel in its airframe with carbon composites could conceivably cut
its
weight to half to one-third of its original dry weight, allowing
higher
altitude or carrying capacity.
*The X-33 though was meant to be only a suborbital test vehicle and
not to
carry payload. However, the image attached below from the "X-33
VentureStar"
page shows it does have a payload bay. This would in fact be large
enough to
carry a two-man cockpit, a la an F-14-style inline seating
arrangement.
*How does the suborbital X-33 attain sufficient delta-v to attain
orbit with
airlaunch? Let's say the Mach 3 carrier craft at X-33 release can give
it an
initial velocity of 1,000 m/s at an altitude of 100,000 ft, 30 km, and
we
wanted to get to orbit at 100 km.
Orbital velocity is about 7,800 m/s. However, we also need to expend
fuel
to get to the desired altitude and to overcome air drag. This page
shows the delta-v
losses due to air drag are surprisingly low:

Drag: Loss in Ascent, Gain in Descent, and What It Means for
Scalability.
Ariane A-44L: Gravity Loss: 1576 m/s Drag Loss: 135 m/s
Atlas I: Gravity Loss: 1395 m/s Drag Loss: 110 m/s
Delta 7925: Gravity Loss: 1150 m/s Drag Loss: 136 m/s
Shuttle: Gravity Loss: 1222 m/s Drag Loss: 107 m/s
Saturn V: Gravity Loss: 1534 m/s Drag Loss: 40 m/s (!!)
Titan IV/Centaur: Gravity Loss: 1442 m/s Drag Loss: 156 m/shttp://gravityloss.wordpress.com/2008/01/10/drag-loss-in-ascent-gain-...

*I'll take the air drag loss in the range of 100 m/s. The gravity loss
depends
on many factors including the altitude of the orbit you want to reach.
A
common way to estimate it is to use the *equation you get from setting
the
potential energy of a certain altitude equal to the kinetic energy of
the
velocity needed to attain that altitude by ballistic launch: v = sqrt
(2gh) .
So at h= 100,000 m, v = sqrt(2*9.8*100,000) = 1,400 m/s. Then the
total
delta-v required would be 7,800 + 1,400 + 100 = 9,300 m/s.
*That 1,400 m/s delta-v just for the altitude of orbit, aside from the
orbital
speed required, is quite large. But key is that this can be reduced if
your
vehicle is given a high initial speed. The idea is that the equation
between
kinetic and gravitational potential energy actually looks like this:
(1/2)Vf^2 -(1/2)vi^2 = 2gh. Then if we have an initial speed of 1,000
m/s and
we only need an additional delta-v to go from our launch altitude of
30 km to
the orbital altitude of 100 km, then the equation becomes:

(1/2)(1540.1)^2 - (1/2)(1000)^2 = 2*9.8*70,000. So this means we need
and
additional delta-v of only 540 m/s to add to our initial 1,000 m/s
speed to
reach the 100 km altitude.
*Then our required total delta-v is 7,800 + 540 + 100 = 8,440. Now to
see how
much delta-v needs to be supplied by the X-33 we'll also include the
fact that
by launching near the equator we get an additional 462 m/s speed for
free, so:
*delta-v still needed = 8,440 - 1,000 - 462 = 6,978 m/s.

*Since the altitude for our launch is so high we can get the full
vacuum Isp
of our engines. Then the mass ratio for a X-33 engine Isp of 455 s and
required delta-v of 6,978 m/s is:

exp(6978/4550) = 4.635, which means the fuel ratio is 3.635.
*Here is where the lightened fuel tanks are needed. The dry mass of
the X-33
was supposed to be 65,000 lb. Then the fuel mass would be 3.635 *
65,000 =
236,275 lb. But the tanks could only hold 210,000 lb of fuel.
*I discussed the proposal for lightening the X-33 tanks he

Passenger market for suborbital, hypersonic transports.http://www.bautforum.com/space-explo...ger-market-sub...

*The idea is that the problem with the X-33 tanks was their unusual,
noncylindrical shape, which *means more tank mass has to be used to
hold the fuel than for usual cylindrical tanks. The suggestion was to
use multiple cylindrical tanks to get the same storage of fuel at a
low weight. This would cut the mass of the two liquid hydrogen tanks
from 9,200 lb. total to 4,600 lb. total, and would cut the mass of the
liquid oxygen tank from 6,000 lb to 1,500 lb, for a total weight
saving of 9,100 lb. Then the dry weight would become 55,900 lb., and
the fuel weight, 203,200 lb, which is within the fuel capacity of the
X-33.

* *Bob Clark

Image of interior of the X-33.http://www.fas.org/spp/guide/usa/lau...3_interior.jpg


II.) The Ares I upper stage including payload was supposed to weigh
about 175,000 kg, 385,000 lb. Two RS-84 kerosene fueled engines in a
flyback booster replacing the solids in the Ares I lower stage would
give a system able to match or exceed the lift capacity of the Ares I.
The X-33 fully fueled only weighed 273,000 lb. Then the kerosene-
fueled flyback booster would be able to give the X-33 a delta-v to
match or exceed that supplied by the Ares I lower stage.
The delta-v supplied by the Ares I lower stage, velocity and altitude,
is about 3,000 m/s. You need about 9,300 m/s delta-v for orbit. Then
the X-33 would have to supply about 6,300 m/s delta-v. But you can get
462 m/s for free from the Earth's rotation near the equator. So let's
call the required delta-v 5,838 m/s. The mass ratio for a delta-v of
5,838 m/s and an engine Isp of 455 sec is: exp(5,838/4550) = 3.61. So
the fuel ratio is 2.61.
The X-33 can hold 210,000 lb of fuel. Then the dry mass including
payload could be 210,000/2.61 = 80,500 lb. The empty weight of the
X-33 was supposed to be 65,000 lb. Then it could carry 15,500 lb
payload either in the small payload bay it has internally or in an
external canister.


Bob Clark
  #3  
Old September 11th 09, 09:45 PM posted to sci.space.policy
Orval Fairbairn[_2_]
external usenet poster
 
Posts: 154
Default The X-33 as the upper stage of a two-stage-to-orbit system.

In article . ath.cx,
noauth wrote:

"Robert Clark" wrote in message news:fcf73abc-
...
I.) In 2006, Aviation Week claimed in an article that the Air Force
had been
testing a two-stage-to-orbit fully reusable launch system:

Two-Stage-to-Orbit ''Blackstar'' System Shelved at Groom Lake?
Mar 5, 2006
By William B. Scott
http://www.aviationweek.com/aw/gener...el=awst&id=new
s/030606p1.xml

TSTO spaceplanes.
http://robotpig.net/aerospace/en_tsto.php


The reason why the U.S. has always kept this system a secret is
probably because it's a violation of the Salt-II FOBS treaty:

"Each Party undertakes not to develop, test, or deploy...)(c)
systems for placing into Earth orbit nuclear weapons or any other kind
of weapons of mass destruction, including fractional orbital missiles"

I'll bet that the SpacePlane had the capability to go into orbit with
nuclear weapons on board. Such a system could deliver a first-strike
within 5 minutes of the order being given if the vehicle was in a
favorable orbital position.

--


WHY would anybody WANT to place nuclear weapons into orbit? That would
make them inaccessible for maintenance, etc. and impose limitations on
accuracy. ICBMs, bombers and SLBMs will perform the strike mission more
simply and are more accurate.

--
Remove _'s from email address to talk to me.
  #4  
Old September 12th 09, 12:08 AM posted to sci.space.policy
Jorge R. Frank
external usenet poster
 
Posts: 2,089
Default The X-33 as the upper stage of a two-stage-to-orbit system.

Orval Fairbairn wrote:
In article . ath.cx,
noauth wrote:

"Robert Clark" wrote in message news:fcf73abc-
...
I.) In 2006, Aviation Week claimed in an article that the Air Force
had been
testing a two-stage-to-orbit fully reusable launch system:

Two-Stage-to-Orbit ''Blackstar'' System Shelved at Groom Lake?
Mar 5, 2006
By William B. Scott
http://www.aviationweek.com/aw/gener...el=awst&id=new
s/030606p1.xml

TSTO spaceplanes.
http://robotpig.net/aerospace/en_tsto.php

The reason why the U.S. has always kept this system a secret is
probably because it's a violation of the Salt-II FOBS treaty:

"Each Party undertakes not to develop, test, or deploy...)(c)
systems for placing into Earth orbit nuclear weapons or any other kind
of weapons of mass destruction, including fractional orbital missiles"

I'll bet that the SpacePlane had the capability to go into orbit with
nuclear weapons on board. Such a system could deliver a first-strike
within 5 minutes of the order being given if the vehicle was in a
favorable orbital position.

--


WHY would anybody WANT to place nuclear weapons into orbit? That would
make them inaccessible for maintenance, etc. and impose limitations on
accuracy. ICBMs, bombers and SLBMs will perform the strike mission more
simply and are more accurate.


Besides, his broad reading of SALT is ludicrous. Under that definition,
the space shuttle and all ELVs would also be banned. Clearly, the space
plane would not be banned simply because it *could* carry nuclear
weapons to orbit.
 




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