Space shuttle for space tourism and first stage of a TSTO.
in [Physics]
|
Prev: Eclipses Yield First Images of Elusive Iron Line in Solar Corona
Next: What happens when all dimensions are doubled overnight?
From: Robert Clark on 25 Jan 2010 11:33 On Jan 15, 5:14 am, Robert Clark <rgregorycl...(a)yahoo.com> wrote: > ... > > Robert Clark <rgregorycl...(a)yahoo.com> wrote: > > > Remember a large mass of propellant carried on the *inside* is the > > >usual way rockets operate. Think of it this way: which would require > > >greater mass and complexity of structural strengthening members, > > > >the S-IC Saturn V first stage carrying the 5,000,000 lbs. of > > >propellant inside, as it actually did, or a huge outside tank hanging > > >off attachments points containing, say, 8,000,000 lbs, with just a big > > >empty space in the rocket between the engines and the second stage? > In regards to its feasibility, a key fact that needs to be kept in mind is how low is the amount of structural strengthening mass required compared to the mass of propellant when the fuel tank is in the *usual position inside the rocket along the center line of the vehicle*. Look at the specifications page on Falcon 9: SpaceX Falcon 9 Updated 4/16/2009 Vehicle Components http://spacelaunchreport.com/falcon9.html#components The propellant mass for the first stage is 279,900 kg, while its dry mass is 14,730 kg. But remember a lot of this dry mass is in the 9 engines of the first stage. This page gives its engine a sea level thrust of 95,000 lbs. at a thrust to weight ratio of 92: SPACEX COMPLETES DEVELOPMENT OF MERLIN REGENERATIVELY COOLED ROCKET ENGINE. https://www.spacex.com/press.php?page=33 So its weight is 1032 lbs, 470 kg. Then 9 weigh 4230 kg. Lox/kerosene propellant tanks also weigh typically 1/100th the propellant weight, so about 2,800 kg. Then less than 7,700 kg is left over that might be devoted to strengthening mass to support this mass of propellant, and additionally to transmit the nearly 1,000,000 lbs of thrust to the rest of the rocket. That is why I'm arguing comparatively little additional strengthening mass is required to begin with according to what has always been the case with inline propellant tanks. And considering that already such strengthening mass has to be present to handle the higher thrust and propellant loads in the orbiter in its present configuration little additional mass to this amount would have to be added, though the placement would likely be different. And in regards to the cost of the addition of our 300,000 kg propellant tank in the orbiter payload bay, the launch cost for Falcon 9 is expected to be ca. $27 to $35 million. This is for the entire two stage rocket with a total of 10 engines. Notice that the development costs are also included in these launch costs. The engines are the hardest part of a rocket development and undoubtedly the cost of these 10 engines represents the lion share of the launch cost. Notice also there must be structural strengthening to connect the nearly 1,000,000 thrust to the rest of the rocket. So this must be part of costs. The conclusion I draw is that it would cost significantly less than $35 million to build and attach the required propellant tank to the shuttle payload bay. Bob Clark
From: Me on 25 Jan 2010 18:51 On Jan 25, 11:33 am, Robert Clark <rgregorycl...(a)yahoo.com> wrote: > The conclusion I draw is that it would cost significantly less than > $35 million to build and attach the required propellant tank to the > shuttle payload bay. > Incorrect conclusion. a. Falcon 9 construction is not applicable to a payload bay propellant tank b. It is impossible to put a 300,000 kg propellant tank in the orbiter payload bay,
From: Robert Clark on 26 Jan 2010 09:33 On Jan 19, 5:56 pm, Robert Clark <rgregorycl...(a)yahoo.com> wrote: > ...This page gives the specifications of the Ares I: > > Space Launch Report - Ares I.http://www.spacelaunchreport.com/ares1.html > > The gross weight including payload is given as 912,660 kg and the > gross weight of the first stage as 732,550 kg. So the gross weight of > the Ares I second stage plus payload is 180,110 kg. > Then the gross weight for the 55,442 kg dry weight of the > reconfigured shuttle, plus 300,000 kg propellant load, plus 180,110 kg > second stage and payload is 535,552 kg, 1,178,214 lbs. But the 3 NK-33 > engines I was suggesting to use only put out a total of 1,018,518 lbs. > of thrust at sea level. For this purpose you would need a fourth > NK-33. The dry weight is now 56,664, the gross weight is 536,774 kg, > 1,180,903 lbs., and the sea level thrust of the 4 engines is 1,358,024 > lbs. > Using the average Isp of the NK-33 as the midpoint of the sea level > and vacuum Isp's at 315 s, the achieved delta-V would be 315*9.8*ln > (536,774/(56,664+180,110)) = 2,527 m/s, comparable to the equivalent > delta-V, speed + altitude, provided by the Ares I first stage. The > achieved delta-V is actually higher than this since the rocket spends > most of the time at high altitude, where the Isp is closer to the > vacuum value. > Note that if you want to increase the delta-V, the space occupied by > the crew compartment is now empty. This gives an additional 74 cubic > meters that could be used for propellant, which amounts to 74,000 kg > additional lox/kerosene propellant that could be carried. > Then we could still use the planned upper stage of the Ares I while > having a significantly lower development cost and per launch cost of > the now reusable first stage. If we only instead wanted suborbital tourism for the vehicle then you would require much less fuel load. Having engine-out capability is a necessary requirement for manned flights. According to this page, the shuttle has a max emergency landing weight of 240,000 lbs, 109,000 kg: Space Transportatin System. http://science.ksc.nasa.gov/shuttle/technology/sts-newsref/sts_overview.html#launch_sites In the calculation in the above post, after removing several subsystems that wouldn't be needed for an unmanned first stage booster I got a 55,442 kg dry weight using three NK-33 engines. However, for suborbital tourism we need the crew seats and environmental systems so I'll add back the 3,250 kg for these to get a dry weight of 58,692 kg. Now assume we max our fuel load for the suborbital tourism use at 50,308 kg, so our max takeoff weight is 109,000 kg, the max allowed for the shuttle for landing under abort modes. Then our delta-V assuming a 315 s average Isp of the NK-33's would be: 315*9.8*ln(109,000/58,692) = 1,911 m/s. This is well above the total equivalent delta-V, speed + altitude, required for reaching the 100 km altitude for space tourism. And the achieved delta-V would actually be higher than this because the actual average Isp is closer to the vacuum value than to the midpoint value. Note that in this configuration we even have 2 engine-out capability since the thrust put out by the NK-33's at sea level is over 300,000 lbs. And since our max weight is at the max allowed for landing the vehicle could even glide to a landing with a full fuel load if all three engines failed as long as we did reach high enough initial velocity for aerodynamic lift to operate. (The orbiter has a respectable lift/drag ratio of 4.5 at subsonic speeds.) Keep in mind that not even jet airliners can land safely during an aborted takeoff if they have all engines out unless they reach sufficient altitude and velocity for lift to operate. This is for the case of just carrying 6 passengers in the crew compartment. The case for carrying a full passenger cabin in the payload bay and fuel only in the wings is a much more complicated analysis and probably not susceptible to an elementary analysis. Bob Clark
From: J. Clarke on 26 Jan 2010 11:04
Robert Clark wrote: > On Jan 19, 5:56 pm, Robert Clark <rgregorycl...(a)yahoo.com> wrote: >> ...This page gives the specifications of the Ares I: >> >> Space Launch Report - Ares >> I.http://www.spacelaunchreport.com/ares1.html >> >> The gross weight including payload is given as 912,660 kg and the >> gross weight of the first stage as 732,550 kg. So the gross weight of >> the Ares I second stage plus payload is 180,110 kg. >> Then the gross weight for the 55,442 kg dry weight of the >> reconfigured shuttle, plus 300,000 kg propellant load, plus 180,110 >> kg second stage and payload is 535,552 kg, 1,178,214 lbs. But the 3 >> NK-33 engines I was suggesting to use only put out a total of >> 1,018,518 lbs. of thrust at sea level. For this purpose you would >> need a fourth >> NK-33. The dry weight is now 56,664, the gross weight is 536,774 kg, >> 1,180,903 lbs., and the sea level thrust of the 4 engines is >> 1,358,024 lbs. >> Using the average Isp of the NK-33 as the midpoint of the sea level >> and vacuum Isp's at 315 s, the achieved delta-V would be 315*9.8*ln >> (536,774/(56,664+180,110)) = 2,527 m/s, comparable to the equivalent >> delta-V, speed + altitude, provided by the Ares I first stage. The >> achieved delta-V is actually higher than this since the rocket spends >> most of the time at high altitude, where the Isp is closer to the >> vacuum value. >> Note that if you want to increase the delta-V, the space occupied by >> the crew compartment is now empty. This gives an additional 74 cubic >> meters that could be used for propellant, which amounts to 74,000 kg >> additional lox/kerosene propellant that could be carried. >> Then we could still use the planned upper stage of the Ares I while >> having a significantly lower development cost and per launch cost of >> the now reusable first stage. > > If we only instead wanted suborbital tourism for the vehicle then you > would require much less fuel load. Having engine-out capability is a > necessary requirement for manned flights. According to this page, the > shuttle has a max emergency landing weight of 240,000 lbs, 109,000 kg: > > Space Transportatin System. > http://science.ksc.nasa.gov/shuttle/technology/sts-newsref/sts_overview.html#launch_sites > > In the calculation in the above post, after removing several > subsystems that wouldn't be needed for an unmanned first stage booster > I got a 55,442 kg dry weight using three NK-33 engines. However, for > suborbital tourism we need the crew seats and environmental systems so > I'll add back the 3,250 kg for these to get a dry weight of 58,692 kg. > Now assume we max our fuel load for the suborbital tourism use at > 50,308 kg, so our max takeoff weight is 109,000 kg, the max allowed > for the shuttle for landing under abort modes. Then our delta-V > assuming a 315 s average Isp of the NK-33's would be: > > 315*9.8*ln(109,000/58,692) = 1,911 m/s. This is well above the total > equivalent delta-V, speed + altitude, required for reaching the 100 km > altitude for space tourism. And the achieved delta-V would actually be > higher than this because the actual average Isp is closer to the > vacuum value than to the midpoint value. > Note that in this configuration we even have 2 engine-out capability > since the thrust put out by the NK-33's at sea level is over 300,000 > lbs. And since our max weight is at the max allowed for landing the > vehicle could even glide to a landing with a full fuel load if all > three engines failed as long as we did reach high enough initial > velocity for aerodynamic lift to operate. (The orbiter has a > respectable lift/drag ratio of 4.5 at subsonic speeds.) > Keep in mind that not even jet airliners can land safely during an > aborted takeoff if they have all engines out unless they reach > sufficient altitude and velocity for lift to operate. If "lift" was not "operating" then an airliner would not have reached _any_ altitude. cretin. <plonk> |