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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 |
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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 |
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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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