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#1




SpaceX BFR tanker as an SSTO.
On Monday, October 23, 2017 at 2:30:22 PM UTC+13, Sylvia Else wrote:
On 23/10/2017 1:37 AM, Robert Clark wrote: The SpaceX BFR tanker can serve as a reusable SSTO by switching to a winged, horizontal landing mode: On a quick scan through, it looks to me as if he's ignored the mass of the landing gear, which in a typical aircraft is 10% of the empty weight. A nonair breathing SSTO ends up with a horribly small payload fraction, which means in turn that only a modest increase in vehicle mass can wipe out the payload completely. Sylvia. With VTOL using rockets, and with an advanced 'active' landing platform, no landing gear is needed. Merely hold down clamps. Catching Balls with Robots https://www.youtube.com/watch?v=R6pPwP3s7s4 Which I discussed here; Active Landing Pads  catch rockets https://vimeo.com/37102557 BFR Mass Budget: BFR system consists of 85 t second stage with 1,250 t propellant capacity  delivering 150 t propellant for transfer to a BFR spaceship  after burning 1,100 t. 8 flights of a freighter refill the 1,200 t capacity BFR spaceship. 85 t / 1250 t = 6.8% structure fraction. Way more than needed. We have 9.75 t of propellant to bring the 85 t structure to a soft landing from 0.34 km/sec terminal velocity. 10% of the 95 t structure + propellant is 9.5 t  set that aside for landing gear. Tank fraction is 3.5% of the original capacity  43.75 t. 7 Raptor engines at 1 t each is 7 t. That's 50.75 t subtotal. 9.5 t landing gear 60.25 t subtotal. Add 9.75 t propellant  and that's 70.00 t. Add 15 t thermal protection and other hardware  and consumables  and you have your 85.00 t budget. * * * https://www.youtube.com/watch?v=tdUX3ypDVwI Specifications Complete................ BFR booster........... BFR spaceship LEO Payload........... reusable: 150 t (330,000 lb) ................................ expendable: 250 t (550,000 lb) Return Payload....... 50 t (110,000 lb) Diameter................. 9 m (30 ft) Length................... 106 m (348 ft)....... 58 m (190 ft) 48 m (157 ft) Maximum mass...... 4,400 t (9,700,000 lb) 1,335 t (2,943,000 lb) Engines.................. 31Ã— SL Raptors 3Ã— SL + 4Ã— vacuum Raptors[37] Thrust.................... 52.7 MN (11,800,000 lbf) 12.7 MN (1,300 tf; 2,900,000 lbf) total Propellant Capacity 1,100 t (2,400,000 lb): 240 t CH4 + 860 t O2 Empty weight......... 85 t (187,000 lb) 
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#2




SpaceX BFR tanker as an SSTO.
With VTOL using rockets, and with an advanced 'active' landing platform, no landing gear is needed. Merely hold down clamps. Catching Balls with Robots https://www.youtube.com/watch?v=R6pPwP3s7s4 Which I discussed here; Active Landing Pads  catch rockets https://vimeo.com/37102557 BFR Mass Budget: BFR system consists of 85 t second stage with 1,250 t propellant capacity  delivering 150 t propellant for transfer to a BFR spaceship  after burning 1,100 t. 8 flights of a freighter refill the 1,200 t capacity BFR spaceship. 85 t / 1250 t = 6.8% structure fraction. Way more than needed. We have 9.75 t of propellant to bring the 85 t structure to a soft landing from 0.34 km/sec terminal velocity. 10% of the 95 t structure + propellant is 9.5 t  set that aside for landing gear. Tank fraction is 3.5% of the original capacity  43.75 t. 7 Raptor engines at 1 t each is 7 t. That's 50.75 t subtotal. 9.5 t landing gear 60.25 t subtotal. Add 9.75 t propellant  and that's 70.00 t. Add 15 t thermal protection and other hardware  and consumables  and you have your 85.00 t budget. Keep in mind there is a difference between the tanker and the spaceship versions of the upper stage. The spaceship version has higher dry mass because it has the passenger quarters and supplies for a 6 month trip to Mars carrying 100 colonists. Since the tanker version doesn't carry this passenger compartment it has much lighter dry mass. This is illustrated by the ITS upper stage introduced last year: https://i.imgur.com/GsyREf7.png The ITS spaceship had a dry mass of 150 tons, and the ITS tanker a dry mass of 90 tons. The halfsize BFR spaceship has a dry mass of 85 tons. Then applying similar scaling to the BFR tanker we can estimate its dry mass as 50 tons. It is the BFR tanker that would be used to deliver payload to LEO. Bob Clark 
#3




SpaceX BFR tanker as an SSTO.
You can buy them and do anything you can get a license to do. So, have at it!
With a 3.2 km/sec exhaust speed and the requirement to boost through a 9.2 km/sec delta vee to attain a 7.91 km/sec orbital speed (with 1.29 km/sec lost due to air drag and gravity losses) we have a propellant fraction of; u = 1  1 / exp(9.2/3.2) = 0.943583860496223 ~ 94.36% So, with a 3.5% structure fraction you could place 2.14% into LEO. With a TSTORLV and the same structure fraction, dividing the delta vee by 2 to make each stage 4.6 km/sec we have; u = 1  1 / exp(4.6/3.2) = 0.762479180904542 ~ 76.25% So with a 3.5% structure fraction you could push 20.25% of each stage through 4.6 km/sec. So, here you have 0.202520819095458 squared which is 0.0410146821670953 ~ 4.1% Nearly double the take off weight into orbit. Let's do three stages! 9.2 km/sec / 3 = 3.067 so, u = 1  1 / exp(3.067/3.2) = 0.616468427123689 ~ 61.65% So with a 3.5% structure fraction you have 0.348531572876311 ~ 34.85% payload on each stage. 0.348531572876311 cubed is 4.23% A slight, but measureable improvement. So, you can see there is a significant benefit to two stages over one. Now, if your structure fraction is higher, the benefit of staging is higher. If we go from 3.5% to 5.0% structure fraction, we have 1 stage to orbit 0.64%  155.85 t per t on orbit 2 stage to orbit  3.51%  28.44 t per t on orbit 3 stage to orbit  3.71%  26.96 t per t on orbit On Sunday, October 29, 2017 at 3:44:43 AM UTC+13, Robert Clark wrote: With VTOL using rockets, and with an advanced 'active' landing platform, no landing gear is needed. Merely hold down clamps. Catching Balls with Robots https://www.youtube.com/watch?v=R6pPwP3s7s4 Which I discussed here; Active Landing Pads  catch rockets https://vimeo.com/37102557 BFR Mass Budget: BFR system consists of 85 t second stage with 1,250 t propellant capacity  delivering 150 t propellant for transfer to a BFR spaceship  after burning 1,100 t. 8 flights of a freighter refill the 1,200 t capacity BFR spaceship. 85 t / 1250 t = 6.8% structure fraction. Way more than needed. We have 9.75 t of propellant to bring the 85 t structure to a soft landing from 0.34 km/sec terminal velocity. 10% of the 95 t structure + propellant is 9.5 t  set that aside for landing gear. Tank fraction is 3.5% of the original capacity  43.75 t. 7 Raptor engines at 1 t each is 7 t. That's 50.75 t subtotal. 9.5 t landing gear 60.25 t subtotal. Add 9.75 t propellant  and that's 70.00 t. Add 15 t thermal protection and other hardware  and consumables  and you have your 85.00 t budget. Keep in mind there is a difference between the tanker and the spaceship versions of the upper stage. The spaceship version has higher dry mass because it has the passenger quarters and supplies for a 6 month trip to Mars carrying 100 colonists. Since the tanker version doesn't carry this passenger compartment it has much lighter dry mass. This is illustrated by the ITS upper stage introduced last year: https://i.imgur.com/GsyREf7.png The ITS spaceship had a dry mass of 150 tons, and the ITS tanker a dry mass of 90 tons. The halfsize BFR spaceship has a dry mass of 85 tons. Then applying similar scaling to the BFR tanker we can estimate its dry mass as 50 tons. It is the BFR tanker that would be used to deliver payload to LEO. Bob Clark 
#4




SpaceX BFR tanker as an SSTO.
"William Mook" wrote in message
... You can buy them and do anything you can get a license to do. So, have at it! With a 3.2 km/sec exhaust speed and the requirement to boost through a 9.2 km/sec delta vee to attain a 7.91 km/sec orbital speed (with 1.29 km/sec lost due to air drag and gravity losses) we have a propellant fraction of; u = 1  1 / exp(9.2/3.2) = 0.943583860496223 ~ 94.36% So, with a 3.5% structure fraction you could place 2.14% into LEO. With a TSTORLV and the same structure fraction, dividing the delta vee by 2 to make each stage 4.6 km/sec we have; u = 1  1 / exp(4.6/3.2) = 0.762479180904542 ~ 76.25% So with a 3.5% structure fraction you could push 20.25% of each stage through 4.6 km/sec. So, here you have 0.202520819095458 squared which is 0.0410146821670953 ~ 4.1% Nearly double the take off weight into orbit. Let's do three stages! 9.2 km/sec / 3 = 3.067 so, u = 1  1 / exp(3.067/3.2) = 0.616468427123689 ~ 61.65% So with a 3.5% structure fraction you have 0.348531572876311 ~ 34.85% payload on each stage. 0.348531572876311 cubed is 4.23% A slight, but measureable improvement. So, you can see there is a significant benefit to two stages over one. Now, if your structure fraction is higher, the benefit of staging is higher. If we go from 3.5% to 5.0% structure fraction, we have 1 stage to orbit 0.64%  155.85 t per t on orbit 2 stage to orbit  3.51%  28.44 t per t on orbit 3 stage to orbit  3.71%  26.96 t per t on orbit Thanks for that calculation. It is correct, for the most part. There is the fact that structure fraction only refers to the rocket stage itself without payload. But when you calculate the fraction that reaches orbit, that's of the entire mass including payload. So it's not quite accurate to subtract off the structure fraction from this. But since the payload is a small percentage of the entire stage mass it's a small discrepancy. However, a key consideration that should be taken into account is that to optimize the payload for a SSTO you really should use altitude compensation. Your structure fraction for the stage of 3.5% is very good, but the exhaust speed of 3.2 km/sec isn't very good for a vehicle you want to be SSTO. Altitude compensation allows you to maximize your vacuum Isp while optimizing your sea level thrust at launch as well. Since we're discussing in the context of the Raptor engine I'll use estimates for methane engines. I'll use a rocket engine analysis program to estimate the possible vacuum Isp with methane fuel: http://www.propulsionanalysis.com/index.htm The specs of the Raptor in its latest incarnation are given he https://www.freelists.org/archives/a...Xl97QZZkJ3.jpg If you use the cited combustion chamber pressure of 250 bar of the Raptor, but give it an expansion area ratio of 300, possible with altitude compensation, then the vacuum Isp can be in the range of 390 s according to the rocket engine analysis program. In contrast, the sea level Raptors to be used on the BFR will have only a vacuum Isp of 356 s, and even the vacuum optimized Raptors to be used only have a vacuum Isp of 375 s. By using altitude compensation you don't have to make these trades of how many sea level vs. vacuum engines to use. All the engines will have the optimal performance both at sea level and at vacuum with altitude compensation. So let's redo your calculation assuming the good structure fraction of 3.5% but using altitude compensation to improve the Isp to 390 s, exhaust speed of 3,824 m/s. Then the proportion of the rocket, dry mass plus payload, that reaches orbit is: 1/exp(9200/3,824) = .09019, then subtracting off the .035 for structure fraction, the payload fraction it would be .09019  .035 = .05519, about 5.5%, significantly better than the 2.14% you get with only a 3.2 km/s exhaust speed. Even more notable is that the payload fraction of the SSTO with altitude compensation is even better than that of the TSTO without it. But to compare apples to apples, let's calculate the payload fraction assuming both stages of the TSTO get the 390 s Isp. In that case, 1/exp(4600/3,824) = .300313 is the final stage mass at burn out, dry mass plus payload. So subtracting off the structure fraction, the payload fraction for each stage would be .300313.035 = .265313. Then squaring this because it is a twostage, the payload fraction would be .07039, about 7%, well above the payload fraction of the TSTO without altitude compensation. This shows altitude compensation can make a significant improvement even for multistage vehicles. The payload fraction is still better than the SSTO case, but it is not as radically better as in the case without altitude compensation. When you take into account you don't have the extra expense of the second stage and of integrating the two stages together, the SSTO can be useful for launching smaller payloads. Bob Clark   Carbon nanotubes can revolutionize 21stcentury technology IF they can be made arbitrarily long while maintaining their strength. Some proposals to accomplish that he From Nanoscale to Macroscale: Applications of Nanotechnology to Production of Bulk UltraStrong Materials. American Journal of Nanomaterials. Vol. 4, No. 2, 2016, pp 3943. doi: 10.12691/ajn422  Research Article. http://pubs.sciepub.com/ajn/4/2/2/  
#5




SpaceX BFR tanker as an SSTO.
SSASTOhttps://en.wikipedia.org/wiki/Douglas_SASSTO
The SII stage of the Saturn V. 5 J2 engines 1,000,000 lbs thrust Take off Weight 1,060,000 lbs weight Structure 80,560 lbs Ve=9,235 mph. This can carry 34,100 lbs into LEO. About the size of a Gemini Capsule! Today, we could do a Dragon Capsule! Total take off weight with payload: 1,094,100 lbs. Now, an aerospike engine, has a higher thrust to weight than a bell nozzle. It also has altitude compensating features which you've described below. https://www.youtube.com/watch?v=0Y0FS8Z1Qk So, if you take 7 J2 pumpsets, and outfit a toroidal aerospike engine around it at the base of the S11, and equip the zero height aerospike engines with a heat shield, you should using modern materials, attain the same structural fraction as was achieved in the 1960s with the SII whilst increasing thrust to 1,400,000 lbs. With a 9,235 mph exhaust speed a thrust of 1,400,000 lbf requires 3,323.8 lbs per second of propellant flow. Acceleration is 42.2 ft/sec/sec minus 32.2 ft/sec/sec is 10.0 ft/sec/sec. If we accelerate straight up for 100 seconds a total of 332,380 lbs of propellant have been burned. At that point the vehicle masses 727,620 lbs. So, acceleration rises at constant thrust to 1,400,000 lbf / 727,620 lbs = 62.2 ft/sec/sec minust 30 ft/sec/sec net that's a = 42.2*exp(0.00388 * t) V = integral(42.2*exp(0.00388*t) 32.2 ) dt, t=0 to 100 = 1,935.68 ft/sec = 1,319.8 mph. D = integral(10,876.3 * exp(0.00388*t)  32.2*t  10876.3) dt, t=0 to 100 = 80,156.1 ft = 15.18 miles Of course, you could start tilting the rocket, and calculate the cosine of the angle as vertical, and the sine of the angle as horizontal add them together to get total velocity, take the ratio to get the tangent and so forth.. av=cos(theta*t)*(42.2*exp(0.00388*t)32.2) ah=sin(theta*t)*(42.2*exp(0.00388*t)) at=sqrt(av^2+ah^2) V=integral(sqrt(av^2+ah^2)) D=integral(integral(sqrt(av^2+ah^2))) At $1,000 per pound, construction cost, the entire system would cost $85 million for the booster $35 million for the payload  $120 million total. Reused 1200 times  the cost is $100,000 per flight. Propellant is $500,000 in quantity and maintenace $82,000  between flights. That's $20 per pound. http://www.spacefuture.com/archive/h..._systems.shtml * * * * On Tuesday, October 31, 2017 at 7:55:12 PM UTC+13, Robert Clark wrote: "William Mook" wrote in message ... You can buy them and do anything you can get a license to do. So, have at it! With a 3.2 km/sec exhaust speed and the requirement to boost through a 9..2 km/sec delta vee to attain a 7.91 km/sec orbital speed (with 1.29 km/sec lost due to air drag and gravity losses) we have a propellant fraction of; u = 1  1 / exp(9.2/3.2) = 0.943583860496223 ~ 94.36% So, with a 3.5% structure fraction you could place 2.14% into LEO. With a TSTORLV and the same structure fraction, dividing the delta vee by 2 to make each stage 4.6 km/sec we have; u = 1  1 / exp(4.6/3.2) = 0.762479180904542 ~ 76.25% So with a 3.5% structure fraction you could push 20.25% of each stage through 4.6 km/sec. So, here you have 0.202520819095458 squared which is 0.0410146821670953 ~ 4.1% Nearly double the take off weight into orbit. Let's do three stages! 9.2 km/sec / 3 = 3.067 so, u = 1  1 / exp(3.067/3.2) = 0.616468427123689 ~ 61.65% So with a 3.5% structure fraction you have 0.348531572876311 ~ 34.85% payload on each stage. 0.348531572876311 cubed is 4.23% A slight, but measureable improvement. So, you can see there is a significant benefit to two stages over one. Now, if your structure fraction is higher, the benefit of staging is higher. If we go from 3.5% to 5.0% structure fraction, we have 1 stage to orbit 0.64%  155.85 t per t on orbit 2 stage to orbit  3.51%  28.44 t per t on orbit 3 stage to orbit  3.71%  26.96 t per t on orbit Thanks for that calculation. It is correct, for the most part. There is the fact that structure fraction only refers to the rocket stage itself without payload. But when you calculate the fraction that reaches orbit, that's of the entire mass including payload. So it's not quite accurate to subtract off the structure fraction from this. But since the payload is a small percentage of the entire stage mass it's a small discrepancy. However, a key consideration that should be taken into account is that to optimize the payload for a SSTO you really should use altitude compensation. Your structure fraction for the stage of 3.5% is very good, but the exhaust speed of 3.2 km/sec isn't very good for a vehicle you want to be SSTO. Altitude compensation allows you to maximize your vacuum Isp while optimizing your sea level thrust at launch as well. Since we're discussing in the context of the Raptor engine I'll use estimates for methane engines. I'll use a rocket engine analysis program to estimate the possible vacuum Isp with methane fuel: http://www.propulsionanalysis.com/index.htm The specs of the Raptor in its latest incarnation are given he https://www.freelists.org/archives/a...Xl97QZZkJ3.jpg If you use the cited combustion chamber pressure of 250 bar of the Raptor, but give it an expansion area ratio of 300, possible with altitude compensation, then the vacuum Isp can be in the range of 390 s according to the rocket engine analysis program. In contrast, the sea level Raptors to be used on the BFR will have only a vacuum Isp of 356 s, and even the vacuum optimized Raptors to be used only have a vacuum Isp of 375 s. By using altitude compensation you don't have to make these trades of how many sea level vs. vacuum engines to use. All the engines will have the optimal performance both at sea level and at vacuum with altitude compensation. So let's redo your calculation assuming the good structure fraction of 3.5% but using altitude compensation to improve the Isp to 390 s, exhaust speed of 3,824 m/s. Then the proportion of the rocket, dry mass plus payload, that reaches orbit is: 1/exp(9200/3,824) = .09019, then subtracting off the .035 for structure fraction, the payload fraction it would be .09019  .035 = .05519, about 5.5%, significantly better than the 2.14% you get with only a 3.2 km/s exhaust speed. Even more notable is that the payload fraction of the SSTO with altitude compensation is even better than that of the TSTO without it. 
#6




SpaceX BFR tanker as an SSTO.
"William Mook" wrote in message
... SSASTOhttps://en.wikipedia.org/wiki/Douglas_SASSTO The SII stage of the Saturn V. Stop, please just stop. If you really want to post this stuff, start your own thread. Please. Stop hijacking other threads.  Greg D. Moore http://greenmountainsoftware.wordpress.com/ CEO QuiCR: Quick, Crowdsourced Responses. http://www.quicr.net IT Disaster Response  https://www.amazon.com/DisasterResp...dp/1484221834/ 
#7




SpaceX BFR tanker as an SSTO.
"Greg \(Strider\) Moore" wrote:
"William Mook" wrote in message ... SSASTOhttps://en.wikipedia.org/wiki/Douglas_SASSTO The SII stage of the Saturn V. Stop, please just stop. If you really want to post this stuff, start your own thread. Please. Stop hijacking other threads. He can't stop Mookjacking threads. I don't think he knows how to start his own.  You are What you do When it counts. 
#8




SpaceX BFR tanker as an SSTO.
Since we're discussing in the context of the Raptor engine I'll use estimates for methane engines. I'll use a rocket engine analysis program to estimate the possible vacuum Isp with methane fuel: http://www.propulsionanalysis.com/index.htm The specs of the Raptor in its latest incarnation are given he https://www.freelists.org/archives/a...Xl97QZZkJ3.jpg If you use the cited combustion chamber pressure of 250 bar of the Raptor, but give it an expansion area ratio of 300, possible with altitude compensation, then the vacuum Isp can be in the range of 390 s according to the rocket engine analysis program. That image of the Raptor on freelists.org site is not available. It's available he http://spaceflight101.com/spx/wpcon...17Musk19.jpg Bob Clark   Carbon nanotubes can revolutionize 21stcentury technology IF they can be made arbitrarily long while maintaining their strength. Some proposals to accomplish that he From Nanoscale to Macroscale: Applications of Nanotechnology to Production of Bulk UltraStrong Materials. American Journal of Nanomaterials. Vol. 4, No. 2, 2016, pp 3943. doi: 10.12691/ajn422  Research Article. http://pubs.sciepub.com/ajn/4/2/2/  
#9




SpaceX BFR tanker as an SSTO.
In article ,
says... Since we're discussing in the context of the Raptor engine I'll use estimates for methane engines. I'll use a rocket engine analysis program to estimate the possible vacuum Isp with methane fuel: http://www.propulsionanalysis.com/index.htm The specs of the Raptor in its latest incarnation are given he https://www.freelists.org/archives/a...Xl97QZZkJ3.jpg If you use the cited combustion chamber pressure of 250 bar of the Raptor, but give it an expansion area ratio of 300, possible with altitude compensation, then the vacuum Isp can be in the range of 390 s according to the rocket engine analysis program. That image of the Raptor on freelists.org site is not available. It's available he http://spaceflight101.com/spx/wpcon...17Musk19.jpg Sure, it's theoretically possible to do such things. But, SpaceX has actually been somewhat conservative on how they've approached engine design. Altitude compensation would add a new facet of "bleeding edge technology" which would introduce more risk into their program (especially schedule risk since this is a relatively unknown area beyond past ground testing). That and they've long ago decided that a fully reusable TSTO is the way to go. If they're successful at landing the BFR first stage on the launch platform, that would eliminate *a lot* of work to get it ready to fly again. They should be able to inspect it, stack the BFR upper stage on top, refuel it, and fly it again. SpaceX hasn't been successful due to using bleeding edge technology. They've been successful by avoiding bleeding edge technology. The one thing they did embrace, because they had to, was supersonic retropropulsion with liquid fueled rocket engines. They only did that because their initial "simpler" plan, which was parachute recovery of first stages in the ocean, quite simply didn't pan out. Jeff  All opinions posted by me on Usenet News are mine, and mine alone. These posts do not reflect the opinions of my family, friends, employer, or any organization that I am a member of. 
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