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Sunday, December 1, 2013

Will the SpaceX push to reusability make Arianespace obsolete?

Copyright 2013 Robert Clark

 By deciding on the solid-fueled Ariane 6, ESA is, unwittingly, betting on SpaceX to fail on reusability. For if SpaceX succeeds then the solid-fueled Ariane 6 becomes obsolete, with billions of dollars and years wasted. ESA would then have to start all over again to develop a liquid-fueled version which can be made reusable:

Musk lays out plans for reusability of the Falcon 9 rocket.
October 3, 2013 by Yves-A. Grondin
Quote:
“The most important thing is that we now believe we have all the pieces of the puzzle (for recovery). If you take the Grasshopper tests, where we were able to do a precision takeoff and landing of a Falcon 9 first stage and you combine it with the results from this flight where we were able to successfully transition from vacuum to hypersonic, through supersonic, through transonic and light the engines all the way through and control the stage all the way through.
“We have all the pieces necessary to achieve a full recovery of the boost stage.”

Falcon 9 first stage in a controlled descent toward the Pacific Ocean. At this point, the stage was about 3 meters (9.8 feet) above the water. (Credit: SpaceX)


  I think it's a bad bet on ESA's part.

 Arianespace has already taken seriously the competition SpaceX offers for their expendable rockets:

SpaceX Challenge Has Arianespace Rethinking Pricing Policies
By Peter B. de Selding | Nov. 25, 2013

Quote:
“I have sent a signal to our customers telling them that I could review our pricing policy, within certain limits,” Israel said in an interview with Les Echos, a French financial newspaper. “I think they have appreciated this.”
Israel’s comments came on the day when Space Exploration Technologies Corp. (SpaceX), after a decade of rattling Arianespace’s cage, is preparing its first-ever launch into the geostationary transfer orbit used by most commercial telecommunications satellites, and the place where most commercial revenue is made.
SpaceX Chairman Elon Musk taunted Arianespace again on Nov. 24, the day before his company’s scheduled launch of the SES-8 satellite owned by SES of Luxembourg.
“Unless the other rocket makers improve their technology rapidly, they will lose significant market share to the Falcon 9,” Musk said in a news briefing.
SpaceX President Gwynne Shotwell added: “Competition is always a good thing. It keeps people sharp. They [Arianespace and other competitors] may not look at it that way, but hopefully they’ll come to appreciate it in the future.”
...
“I am looking at our pricing policy and if we must adapt it to the competition, we will,” Israel said. “We’ll look at the overall efficiency of the Ariane business with a view to optimizing it.”
Israel said there are more small telecommunications satellites being designed now than ever, a fact he attributed in part to the arrival of SpaceX, which has stimulated the market.
http://www.spacenews.com/article/launch-report/38331spacex-challenge-has-arianespace-rethinking-pricing-policies
 IF SpaceX succeeds in cutting prices by reusability, then no readjustment of the pricing will be effective. SpaceX is already undercutting them on pricing and if reusability really does cut the SpaceX prices again by a factor of 4 to 10 then ArianeSpace simply will not be able to compete.

 This will be all due to ESA's decision to go backwards in technology and not forwards in selecting a solid-fueled version of the Ariane 6. Every other space agency in the world will be able to adapt their liquid fueled rockets to make them reusable to match SpaceX's pricing. Only ESA will be left behind - both technically and economically.

 This becomes really bad because they will no longer have the smaller satellites to partially pay for the Ariane 5 launches. This could mean they also lose their entire Ariane 5 market as well! Their entire market for any of their launches will be gone all due to the choice to move backwards in technology.

 Ironically, this would mean their real reason for selecting the solid-fuel Ariane 6 would have no meaning as well. The actual reason why France and Italy want the solid-fueled Ariane 6 is to help defray the costs of the solid-fueled ballistic missiles of the French military and the solid-fueled Vega rocket largely built in Italy. But if SpaceX succeeds in cutting costs by reusability then neither the solid-fueled Ariane 6 nor the Vega, will be used because they will be priced far outside the market. So neither of them will wind up defraying the costs of other solid-fueled rockets in Europe anyway.

 Interestingly, IF SpaceX succeeds in their next test of reusability in Feb. 2014, this might provide an incentive for ESA to at least "hedge their bets" and engage in some development research of adding a second Vulcain to the Ariane 5 core. Then they would not be years behind the other space agencies in the world IF SpaceX succeeds in cutting costs by reusability. 


  Bob Clark

Thursday, November 7, 2013

The Coming SSTO's: Falcon 9 v1.1 first stage as SSTO, Page 2.

Copyright 2013 Robert Clark

 Though in the first test flight of the new version of the Falcon 9, the F9 v1.1, they did not stably "land" the first stage, SpaceX is optimistic they can solve the problem to get a reusable first stage:

SpaceX Hit Huge Reusable Rocket Milestone with Falcon 9 Test Flight (Video).
By Mike Wall, Senior Writer   |   October 17, 2013 02:01pm ET
Combining information from the Falcon 9 v1.1's maiden flight and the ongoing Grasshopper tests should help bring a rapidly reusable rocket closer to reality, SpaceX officials said.
"SpaceX recovered portions of the [Falcon 9 v1.1's first] stage and now, along with the Grasshopper tests, we believe we have all the pieces to achieve a full recovery of the boost stage," they wrote in the Oct. 14 update.
http://www.space.com/23230-spacex-falcon9-reusable-rocket-milestone.html

 SpaceX also plans to transition the half-scale Grasshopper VTVL test vehicle to a full scale Falcon 9 first stage:

Final flight of Grasshopper v1.0 sets new record.
By Brian Dodson
October 14, 2013
http://www.gizmag.com/grasshopper-retires-altitude-record/29384/

 This article says this "Grasshopper 2", as it were, would have all 9 engines of the regular F9 first stage. However, discussions on other forums have said it would only have 3 engines. That would make sense since on stage return, you are using at most 3 engines, and moreover this way, you would not be risking an expensive 9 copies of the Merlins during these Grasshopper test flights.

 Still, in point of fact there would be an advantage of using all 9 engines on this first stage Grasshopper, and with a full propellant load. In November, 2012 Elon Musk gave a lecture in London at the Royal Aeronautical Society.




 About 30 minutes in, he gave the propellant fraction of the new Falcon 9 v1.1 first stage as around 96%, or perhaps 95.5%. The 96% propellant fraction number gives a 25 to 1 mass ratio. But at an Isp of 311 s for the Merlin 1D, the rocket equation gives a delta-v of 311*9.81ln(25) = 9,800 m/s. Since the delta-v to orbit is only about 9,100 m/s, this would allow a significant amount of payload. Then using the 9 engines and the full propellant load on the F9 first stage would allow in fact not just a VTVL test vehicle, but in fact a fully reusable and fully orbital vehicle.  Amusingly, about 36 minutes into Elon's lecture someone asks a question about what he sees as the next big breakthrough in rockets after full reusability. Elon thinks for awhile and can't come up with an answer. He finally jokes maybe warp drive. Ironically, he already has the next big advance: a reusable SSTO.
 
In the blog post "The Coming SSTO's: Falcon 9 v1.1 first stage as SSTO" I had already discussed the F9 v1.1 first stage being used as a SSTO. But there I actually used the side boosters of the Falcon Heavy, which are based on the F9 v1.1 first stage, since they were supposed to have such a high mass ratio, at 30 to 1. However, this information in Elon's lecture on the first stage of the F9 v1.1 suggests it itself would have a surprisingly high mass ratio.
 We'll enter this data into Dr. John Schilling's launch performance calculator to estimate the payload it could carry. On the SpaceX page on the Falcon 9 v1.1 the vacuum thrust is given as 6,672 kN. The Merlin 1D has a vacuum Isp of 311 s. We need to know the propellant mass of the F9 v1.1 first stage.

  I found the propellant loads for the Falcon 9 v1.1 first stage and second stage in this environmental impact report: 

Draft Environmental Impact Statement: SpaceX Texas Launch Site. 
http://www.faa.gov/about/office_org/headquarters_offices/ast/environmental/nepa_docs/review/documents_progress/spacex_texas_launch_site_environmental_impact_statement/media/SpaceX_Texas_Launch_Site_Draft_EIS_V2.pdf  

  They're given on page 66, by the PDF file page numbering:

First and Second Stages  
The first stage of the Falcon 9 is approximately 12 ft by 150 ft, and includes nine Merlin 1D engines. The Merlin engine produces 146,000 lbs of thrust and contains a pump-fed gas generator cycle, turbine exhaust roll control, and hydraulic thrust-vector control. The first stage consists of aluminum LOX and RP-1 tanks that hold approximately 62,000 gallons (gal) of LOX and 38,000 gal of RP-1. The second stage is approximately 12 ft by 41 ft, not including the fairing and payload, and uses one Merlin vacuum engine. The fairing (the top portion of the vehicle where the payload is encapsulated) would be 17 ft by 35 ft, and a smaller version may also be used. The second stage consists of approximately 15,000 gal of LOX and 9,000 gal of RP-1 in tanks with a common bulk head.  

 The conversion factor between gallons and liters is 1 gallon to 3.7854 liters. So the amount of LOX in liters is 62,000*3.7854 = 234,700 liters, 234.70 m³. And the amount in liters of RP-1 is 38,000*3.7854 = 143,800 liters, 143.8 m³. The density of LOX is 1140 kg/m³ and the density of RP-1 is 820 kg/m³. So the total mass of propellant is 1140*234.7 + 820*143.8 = 385,500 kg. This is in the range of what has been estimated for instance on Ed Kyle's site of ca. 389 metric tons (mT): 

Space Launch Report:  SpaceX Falcon 9 v1.1 Data Sheet. 
http://spacelaunchreport.com/falcon9v1-1.html#components  

 However, there is a significant difference from the estimates and the actual propellant load of the second stage. Using the same conversion factors and densities, the total propellant for the second stage is 92,670 kg, whereas Kyle gives it as approx. 64 mT.  

 Take, optimistically, the propellant fraction of the F9 v1.1 first stage as 96%, corresponding to a mass ratio of 25 to 1. Then the dry mass is 16,040 kg, which I'll round to 16,000 kg.  In the calculator input the vacuum values for the Isp and thrust, as it takes into account the reduction at sea level. Select "No" for the "Restartable Upper Stage" option, otherwise the payload will be reduced. Select Cape Canaveral as the launch site at a launch inclination of 28.5 degrees to match the latitude of the launch site. Then the calculator gives an estimated payload of 5,147 kg: 

Mission Performance:
Launch Vehicle:   User-Defined Launch Vehicle
Launch Site:   Cape Canaveral / KSC
Destination Orbit:  185 x 185 km, 28 deg
Estimated Payload:   5147 kg
95% Confidence Interval: 1242 - 9908 kg

 This is surprisingly high for a stage using engines without an especially high Isp. However an SSTO reaches its best performance when using altitude compensation. Let us suppose we use altitude compensation so that the engines on the first stage have the same vacuum Isp as the Merlin Vacuum at 340 s. 
 Note that because of the higher Isp, the thrust is also increased. On that SpaceX page on the Falcon 9 v1.1, the thrust of the single Merlin Vacuum on the upper stage is given as 801 kN. So 9 would have a thrust of 7209 kN, which I'll round to 7,210 kN. Select "Optimal" in the calculator for the "Trajectory". Then the calculator gives the result:

Mission Performance:
Launch Vehicle:   User-Defined Launch Vehicle
Launch Site:   Cape Canaveral / KSC
Destination Orbit:   185 x 185 km, 28 deg
Estimated Payload:   12068 kg
95% Confidence Interval:   7319 - 17788 kg
 This is remarkable as being near the payload cited by SpaceX for the full two stage Falcon 9 v1.1 of 13,150 kg.  

 But for a fair comparison we should see also how high the payload would get for the two stage F9 when altitude compensation is also given to the first stage. The calculation here is made difficult by the fact that we don't know the propellant fraction of the upper stage, so we can't calculate the dry mass from the known propellant mass of 92,670 kg.
 For the upper stage much smaller than the first stage, the mass ratio would not be as great. It is known that as you scale up a rocket the mass ratio improves. The reverse is also true, when you scale down a stage the mass ratio becomes worse. The acceleration at burn out for just an empty upper stage, and payload would also be rather high. Then I'll take the mass ratio for the upper stage at only 10 to 1, giving a 9,200 kg upper stage dry mass. Let's calculate first what the calculator gives as the payload for the present case using the standard Merlin 1D at 311 s Isp. The calculator gives:
Mission Performance:
Launch Vehicle:   User-Defined Launch Vehicle
Launch Site:   Cape Canaveral / KSC
Destination Orbit:  185 x 185 km, 28 deg
Estimated Payload:   13831 kg
95% Confidence Interval: 10061 - 18407 kg
 Rather close to the actual value of 13,150 kg. Now we'll calculate it for the case where the first stage has been given altitude compensation to get a 340 s Isp. We'll change the Isp input to 340 s and also increase the thrust to 7,210 kN as before. Then the calculator gives:


Mission Performance:
Launch Vehicle:   User-Defined Launch Vehicle
Launch Site:   Cape Canaveral / KSC
Destination Orbit:  185 x 185 km, 28 deg
Estimated Payload:   17056 kg
95% Confidence Interval: 12781 - 22223 kg
 This is a significant increase but not nearly as dramatic as the increase for the SSTO case. For the SSTO case the payload more than doubled. But for the TSTO case it increased by less than 25%.

 This could mean the SSTO could approach that of the TSTO on a cost per kilo basis. Elon Musk has said the Falcon 9 first stage takes up about three-quarters of the cost of the Falcon 9:
Musk lays out plans for reusability of the Falcon 9 rocket
October 3, 2013 by Yves-A. Grondin 
Performance hit for reusable rockets:
Musk also addressed the performance hit that results from reserving propellant for landing the first stage.
“If we do an ocean landing (for testing purposes), the performance hit is actually quite small, maybe in the order of 15 percent. If we do a return to launch site landing, it’s probably double that, it’s more like a 30 percent hit (i.e., 30 percent of payload lost).”
...
Musk believes that the most revolutionary aspect of the new Falcon 9 is the potential reuse of the first stage “which is almost three-quarters of the cost of the rocket.”

http://www.nasaspaceflight.com/2013/10/musk-plans-reusability-falcon-9-rocket/
 This would put it at about $40 million out of the $54 million for the full rocket. Then the cost per kilo for the SSTO would be $40,000,000/12,068 = $3,314 per kilo, while for the TSTO it would be $54,000,000/17,056 kg = $3,166 per kilo.

 The benefits of the SSTO would be even more dramatic in the reusable case. In the Nasaspaceflight.com article Elon says the loss in payload for the F9 for returning just the first stage to the launch site was about 30%. This is interesting because he said in another interview the loss in payload for returning both stages would be a loss of about 40%:

Elon Musk on SpaceX’s Reusable Rocket Plans.
By Rand Simberg
February 7, 2012 6:00 PM

Despite the dangers, Musk is clearly a fan of the rocket-powered approach. He told PM that SpaceX has come up with a solution to make both the lower and upper stages of the Falcon 9 reusable. (The Dragon capsule that will fly atop the rocket has already demonstrated that it can be recovered in the ocean after it splash-lands with a parachute, though SpaceX is building vertical-landing capability into that as well.)
The key, at least for the first stage, is the difference in speed. "It really comes down to what the staging Mach number would be," Musk says, referencing the speed the rocket would be traveling at separation. "For an expendable Falcon 9 rocket, that is around Mach 10. For a reusable Falcon 9, it is around Mach 6, depending on the mission." For the reusable version, the rocket must be traveling at a slower speed at separation because the burn must end early, preserving enough propellant to let the rocket fly back and land vertically. This also makes recovery easier because entry velocities are slower.
However, the slower speed also means that the upper stage of the Falcon rocket must supply more of the velocity needed to get to orbit, and that significantly reduces how much payload the rocket can lift into orbit. "The payload penalty for full and fast reusability versus an expendable version is roughly 40 percent," Musk says. "[But] propellant cost is less than 0.4 percent of the total flight cost. Even taking into account the payload reduction for reusability, the improvement is therefore theoretically over a hundred times."

http://www.popularmechanics.com/science/space/rockets/elon-musk-on-spacexs-reusable-rocket-plans-6653023 


 These two quotes together could mean the payload loss from making the upper stage also reusable is 10%, assuming Elon was being consistent between the two quotes. Then a question arise: would the payload loss from the making the SSTO reusable also be just 10% of the payload? 

 This doesn't seem likely, for if you changed the relative sizes of the first and upper stages while keeping the payload the same, then the extra added components for the upper stage such as heat shield, landing legs, and propellant reserve for landing should also change. It should not stay as the same 10% of the payload, regardless of the size of the stage. So we'll need to do use some other sources to see how much payload would likely be lost under the reusable SSTO case.

Payload Lost for a Reusable SSTO.

 We need a heat shield, landing legs, and reserve propellant for the landing. This interesting discussion between noted space-historian Henry Spencer and a former manager for both the DC-X and X-33 programs, Mitchell Burnside Clapp, is about the relative benefits of horizontal versus vertical landing of RLV's:

Horizontal vs. vertical landing (Henry Spencer; Mitchell Burnside Clapp).
http://yarchive.net/space/launchers/horizontal_vs_vertical_landing.html 

 Burnside Clapp conservatively estimates the propellant that needs to be kept on reserve for the landing amounts to about 30 seconds of engine firing. Spencer optimistically estimates it might be as low as 10 seconds. I'll estimate it as 20 seconds. Assume the engine used for the landing has similar sea level Isp as the Merlin at 282 s. But this is not for the full firing of all engines as would be needed for takeoff of a fully loaded rocket. 

 We'll assume we only need enough thrust for the dry mass of the stage, as the needed reserve propellant is a small proportion of this. Taking the dry mass of the first stage as 16,000 kg, 157,000 N, the flow rate of such an engine would be (flow rate) = (thrust)/(exhaust velocity) = 157,000N/2370m/s = 57.5 kg/s. And the propellant for a 20 second burn would be 1,150 kg, 7% of dry mass.

 For the heat shield, it will be the PICA-X material of SpaceX. The mass for this heat shield  used for the Dragon has been estimated in the range of 226 kg. However, the video SpaceX has released of a reusable Falcon 9 shows a heat shield on the upper stage that extends partially down the side of the stage. Then I'll estimate the mass as double that of the Dragon at 550 kg.

 For the landing gear the example of the lighweight gear for the B-58 suggests it can be as low as 1.5% of the landing weight:

Landing gear weight (Gary Hudson; George Herbert; Henry Spencer).
http://yarchive.net/space/launchers/landing_gear_weight.html 


 With lightweight composites this might be reduced to 1% of the landed weight, 160 kg. The total of all three of these extra systems for reusability would then be 1,860 kg, about 12% of the 16,000 kg dry weight. 

 This would need to be subtracted off from the delivered mass to LEO. Then the reusable F9 v1.1 first stage would have a payload to LEO of 10,200 kg.

Comparsion of Costs of Reusable SSTO, Partially Reusable TSTO, and Fully Reusable TSTO.

  First, under the partially reusable case of just the first stage being reusable, this would subtract off 30% of the payload, so from 17,056 kg to 11,940 kg. Now assume the first stage is reusable 10 times and this cuts the cost of that stage by a factor of 10, so to $4 million per flight. Then the upper stage being expendable would be $14 million, i.e. $54 million - $40 million, and the total cost would be $18 million per flight, at a cost per kilo of $1,500 per kilo.

 Now compare to the reusable SSTO case. Again assume 10 uses at a cost of $4 million per flight. Use the reusability loss estimate above that lowers the payload to LEO to 10,200 kg. Then the cost per kilo would be only $390 per kilo(!)

 Perhaps a fairer comparison though would be to the fully reusable TSTO case. This would cut the payload by 40% so from 17,056 kg to 10,230 kg. Since we're using the full rocket 10 times, assume the cost is cut to $5.4 million per flight. This would be a cost per kilo of $527 per kilo. So the reusable SSTO would carry about the same payload but at a better cost per kilo.

 Admittedly though this conclusion is based on very rough estimates for the propellant reserve needed for landing and the mass needed for the heat shield for a long rocket stage compared to that of a capsule.


   Bob Clark


Update, October 18, 2014:

 The calculations here were assuming the Falcon 9 v1.1 had payload to LEO of 13,150 kg. However, as discussed in the post "Golden Spike" Circumlunar Fights, Page 2 this payload is actually that of the partially reusable version. The actual payload of the expendable version is ca. 16,600 kg. 

Then assuming altitude compensation increases the payload of a TSTO by 25%, the Falcon 9 v1.1 with altitude compensation on the first stage would have a payload of ca. 20,000 kg. So in the last section with comparisons of the price per kilo of a reusable SSTO and TSTO, the fully reusable TSTO with 40% loss should have a payload of 12,000 kg. This would still mean the reusable SSTO would have a lower price per kilo than the fully reusable TSTO.


UPDATE, October 25, 2014:

 SSTO's achieve their best usefulness with altitude compensation. Low cost methods of giving already existing engines altitude compensation are discussed here:

Altitude compensation attachments for standard rocket engines, and applications.
http://exoscientist.blogspot.com/2014/10/altitude-compensation-attachments-for.html

 

Thursday, October 31, 2013

A SpaceX Heavy Lift Methane Rocket.

Copyright 2013 Robert Clark

 SpaceX has announced development of a new 300 metric ton (mT), 660,000 lb, thrust engine, the Raptor:

SpaceX Could Begin Testing Methane-fueled Engine at Stennis Next Year.
By Dan Leone | Oct. 25, 2013

 This is supposed to be used for a proposed heavy lift rocket to be used for manned Mars missions. However, I'm not a fan of the 9 engine arrangement used on the Falcon 9, and even less so of the 27 engines proposed for the Falcon Heavy. I would hope that SpaceX would transition to the larger engines for these rockets as well.

 We can do an estimate of the size and payload capacity of the methane-fueled heavy lift rocket. Previous statements from SpaceX have suggested the core of the rocket might be 7 meters wide. However, I wanted to use an 8 meter wide core to make use of the tooling used for the shuttle external tank to save on costs. If we used the same size tank as the shuttle ET then we can calculate the mass of propellant could be carried as methane-lox instead of hydrogen-lox by comparing their densities.

SSTO Case.

 This report by Dr. Bruce Dunn gives densities and performance data on several propellant combinations:

Alternate Propellants for SSTO Launchers.
Dr. Bruce Dunn
Adapted from a Presentation at:
Space Access 96
Phoenix Arizona
April 25 - 27, 1996
https://web.archive.org/web/20140215015634/http://www.dunnspace.com/alternate_ssto_propellants.htm

 In Table 1 the density of methane-lox is 828 kg/m^3 and for hydrogen-lox, 358 kg/m^3. So the same volume would hold 2.4 times more methane-lox. This would put it in the range of 1,700 mT for the methane-lox. Actually it would probably be a little more than this because likely SpaceX would use common bulkhead design for the tank which would mean it could hold more propellant.

 There have been some estimates proposed for this launcher that use 7 copies of Raptor engine on the core. This many probably would be needed when you take into account the reduction in thrust at sea level if using a 1,700 mT sized tank. However, I wanted to keep the maximum number of engines on a core to be at most what was used on the Saturn V at 5 engines. Therefore I'll reduce the propellant load to 1,000 mT.

 For the dry mass, note that Elon has said that the Falcon 9 v1.1 first stage has a propellant fraction in the range of 96%, for a mass ratio of 25 to 1. As you can see in Dunn's Table 1 the density of methane-lox is about 80% that of kerosene-lox. So I'll estimate the mass ratio for the core as 20 to 1. This will put the dry mass of the core at 52,630 kg, which I'll round off to 50,000 kg.

 The vacuum thrust in kilonewtons for 5 Raptors will be 5*300*9.81 = 14,715 kN. We'll calculate the payload for this core stage first as an SSTO. Input these numbers into Dr. John Schillings Launch Performance Calculator. Select "No" for the "Restartable Upper Stage" option, otherwise the payload will be reduced. Use the default altitude of 185 km. Select Cape Canaveral as the launch site with an orbital inclination of  28.5 degrees to match the latitude of the launch site. Then the Calculator gives the result:

Mission Performance:
Launch Vehicle:  User-Defined Launch Vehicle
Launch Site:  Cape Canaveral / KSC
Destination Orbit:  185 x 185 km, 28 deg
Estimated Payload:  42487 kg
95% Confidence Interval:  28319 - 59338 kg

 Two stage case.  

 For the two stage case, I'll take the the upper stage as using a single Raptor and at 1/5th the size of the first stage, so at 200 mT propellant mass and 10 mT dry mass. Enter in 2,943 kN for the thrust of a single Raptor in the column for the second stage and select "Optimal" for the trajectory. Then the Calculator gives:

Mission Performance:
Launch Vehicle:  User-Defined Launch Vehicle
Launch Site:  Cape Canaveral / KSC
Destination Orbit:  185 x 185 km, 28 deg
Estimated Payload:  77569 kg
95% Confidence Interval:  64244 - 93424 kg

  However, for this upper stage likely you won't be able to get as good a mass ratio as the first stage since it would undergo a higher acceleration as the propellant is burned off. This would require a stronger and therefore heavier structure. Then the payload would be reduced below this, though likely still above ca. 70 mT.

Cross-Feed Fueling for Multiple Cores.

 For higher payloads we'll use a combination of 2 or 3 cores. For both of these we'll use cross-feed fueling. To emulate cross-feed fueling with the Schilling Calculator, note that during the parallel burn portion of the flight the propellant for the center core engines is coming from the side booster stage(s). This ensures that the center core will have a full propellant load during its solo burn portion of the flight, after the side booster(s) are jettisoned. 

 So the total amount of propellant burned during the parallel burn portion, is that of the side booster(s) only. But the Schilling Calculator assumes the amount of propellant burned in the center core during the parallel burn is the same as the amount burned in each side booster. So enter in the Calculator for the booster propellant load a fraction of the actual propellant load of a core equal to the number of side boosters divided by the number of cores. So if you're using 2 cores with one used as a side booster enter in the Calculator booster column 1/2 the amount of the actual core propellant load. And if using 3 cores with 2 used as side boosters, enter in 2/3rds the actual core propellant load in the booster section. This will ensure the Calculator interprets the total propellant burned during the parallel burn portion is that of the actual side booster(s) only.

 But you also want the Calculator to take the amount of propellant burned during the center core's solo burn portion of the flight as that of a full propellant load. Since it is already taking it to have burned the same amount as what the side boosters have burned during the parallel burn portion, add this amount onto the actual propellant load of a core and enter this into a first stage column of the Calculator. For the other specifications for both booster(s) and center core such as Isp, dry mass, and thrust enter in the actual values.

2 Core Version.
 Here one core will be used as a side booster. As described above, to emulate cross-feed fueling enter in the Calculator only 500,000 for the propellant load of the booster. Enter in though the actual dry mass of 50,000 kg, actual thrust of 14,715 kN, and actual Isp of 380 s. And for the center core, enter in the first stage column for the propellant 1,500,000 kg, but the real dry mass, thrust, and Isp values. Also use all the actual values for the second stage. Then the Calculator gives the result:

Mission Performance:
Launch Vehicle:  User-Defined Launch Vehicle
Launch Site:  Cape Canaveral / KSC
Destination Orbit:  185 x 185 km, 28 deg
Estimated Payload:  145339 kg
95% Confidence Interval:  121566 - 173707 kg
 This is surprisingly high. However, another consideration besides the fact that the second stage mass ratio likely won't be as good as used here, is that as propellant is burned off during the parallel burn portion, the engines will have to be gimbaled because the propellant is only coming from the side booster stage. This will reduce the payload somewhat.

3 Core Version.
 Here two cores will be used as side boosters. As discussed, to emulate cross-feed we'll enter in the booster column for the propellant load, 2/3rds the actual amount, so only 660,000 kg. And for the center core's propellant load, enter into the first stage column 1,660,000 kg. All the other specifications are given their actual values. Then the Calculator gives:

Mission Performance:
Launch Vehicle:  User-Defined Launch Vehicle
Launch Site:  Cape Canaveral / KSC
Destination Orbit:  185 x 185 km, 28 deg
Estimated Payload:  211268 kg
95% Confidence Interval:  177052 - 251940 kg

  Remarkably high. Twice the payload of the SLS at about the same gross mass.

   Bob Clark

UPDATE, May 1, 2015:

 To get such high performance you would need the lower stage engines to have the high vacuum Isp of 380 s. But lower stage engines usually compromise their performance to use a single nozzle that can work both at low altitudes and high altitudes. Ideal then would be a nozzle that could adapt to the altitude, an altitude compensating engine.
 Some possibilities for this are discussed here:

Altitude compensation attachments for standard rocket engines, and applications.
http://exoscientist.blogspot.com/2014/10/altitude-compensation-attachments-for.html

Saturday, October 5, 2013

DARPA's Spaceplane: an X-33 version.

Copyright 2013 Robert Clark



 DARPA has announced that it will be funding research into a reusable first stage booster to carry an orbital upper stage. But looking at the specifications of the cancelled programs the DC-X's suborbital follow-on, the DC-X2, and on the X-33 you'll note that they each could have performed this role. This would have led to greatly reduced orbital costs. Then both programs were cancelled prematurely.

 Part of the problem is that they were viewed as purely demonstration or experimental programs, without any potential profitability of their own. The profitability would have come with the full, and expensive, SSTO programs to follow. However, if it had been noted these could have been used as fully reusuable first stages, then their value would have been seen on their own. So that they would have been understood as deserving of funding whether or not the SSTO's were to follow.

 The story of the X-33 is well-known now among space advocates:

X-33/VentureStar – What really happened.
January 4, 2006 by Chris Bergin
http://www.nasaspaceflight.com/2006/01/x-33venturestar-what-really-happened/

 It was to be a suborbital experimental test vehicle for a larger SSTO called the VentureStar. For the VentureStar to have been SSTO with significant payload would have required aggressive weight saving techniques such as composite tanks. Such composite tanks were to be tested on the X-33 before committing to the full VentureStar.

 However, the composite tanks failed on the X-33. Since it was felt the SSTO version could not succeed with regular metal tanks, the program was cancelled. However, in point of fact even if you replaced the failed composite tanks with aluminum-lithium ones the X-33 could still be used as a reusable first stage.

 The problem with the tanks is that their unusual conformal shape required them to use greater tank mass compared to the mass of propellant carried than by usual cylindrically shaped tanks:

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 McDonnell-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

  However, ironically it turned out that the hydrogen tank weight for the X-33 actually went down when replaced by aluminum:

From "X-33/VentureStar – What really happened" :
Faced with a project failure, Lockheed Martin and X-33 NASA managers gave the green light to proceed with the fabrication of the new tank. Ironically this new tank weighed in less than the composite tank – disproving one of the reasons for going with a composite tank in the first place.
While the aluminium LH2 tank was much heavier than the composite tank in the skins, the joints were much lighter, which was where all the weight in the composite tank was, due to the multi-lobed shape of the tank requiring a large amount of surrounding structure, such as the joints. Ironically, the original design of the X-33 on the drawing board had the tanks made out of aluminium for this reason – but the cost played a factor for the potential customer base.
Then on replacing the composite hydrogen tanks with Al-Li the dry mass should be less. So I'll use the same numbers for the dry mass and gross mass, 75,000 lbs for the dry mass and 285,000 lbs for the gross.

 The X-33 was to use two aerospike XRS-2200 engines. According to Wikipedia, the XRS-2200 produces 204,420 lbf (909,300 N) thrust with an Isp of 339 seconds at sea level, and 266,230 lbf (1,184,300 N) thrust with an Isp of 436.5 seconds in a vacuum. So two will have a vacuum thrust of 2,368,600 N.

 Now choose for the upper stage an efficient cryogenic stage such as the Centaur or the Ariane H10. We'll use Dr. John Schilling's Launch Performance Calculator to estimate the payload possible. Take the specifications for the Centaur rounded off as 2,000 kg dry mass, 21,000 kg propellant mass, 100 kN vacuum thrust and 451 s vacuum Isp. Then the Calculator gives a payload of 5,275 kg to orbit:

Mission Performance:
Launch Vehicle:  User-Defined Launch Vehicle
Launch Site:  Cape Canaveral / KSC
Destination Orbit:  185 x 185 km, 28 deg
Estimated Payload:  5275 kg
95% Confidence Interval:  4252 - 6507 kg

 The cost of a Centaur upper stage is in the range of $30 million. But how much for a reusable X-33? This article gives the cost to build a X-33 as $360 million in 1998 dollars:

Adventure star  
12:00 18 Nov 1998  Source:  Flight.
By:  Graham Warwick/WASHINGTON DC

 Even taking into account inflation the cost should not be terribly much more than that when you also take into account the decrease in price for composites because of their more common use. 

 The launch preparation costs should also be low since the X-33 was expected to be operated by only a 50 man ground crew compared to the 18,000 required for the shuttle system:

Lockheed Secret Projects: Inside the Skunk Works.

 Say the builder expected a 25% profit over cost of the vehicle over 100 flights. That would be a charge of $4.5 million per flight. With the Centaur upper stage that would be $34.5 million per flight for 5,275 kg to orbit, about $6,500 per kilo. This is a significant saving over the ca. $10,000 per kilo for launchers in the West. It is still well above DARPA's desired price point of $5 million per flight, but it is for a larger payload than the DARPA required 3,000 to 5,000 pounds.

 A lower cost launcher could be obtained using a cheaper upper stage, such as the Ariane H10 stage. This is about 12 mT in propellant load and 1.2 mT in dry mass at 445 s vacuum  Isp and 63 kN vacuum thrust. The Calculator gives a payload mass of 3,762 kg.

 The cost for the H10 stage according to Astronautix is $12 million. Then the total would be $16.5 million. At a payload of 3,672 kg, this is $4,500 per kilo. This would be a great cut in cost for small size payloads, but the total cost is still too high for the DARPA price requirements.

 Another possibility for a cheaper upper stage would be the Falcon 1's first stage. This has a dry mass of 1,450 kg and propellant mass of 27,100. We'll use for it though the upgraded Merlin 1D Vacuum at 800 kN vacuum thrust and 340 s Isp. Then the Calculator gives a payload mass of 5,238 kg. 

 The latest listed price for the Falcon 1 in 2008 was about $8 million. But we only need the first stage. Elon Musk has said for the Falcon 9 the cost of the first stage is 3/4ths the cost. If also true for the Falcon 1, that would put the cost at $6 million for the first stage. Then the total cost would be $10.5 million, $2,000 per kilo. This is a quite low cost per kilo and it would be a significant advance to have payload this size launched at such low cost, whether or not it would qualify under the DARPA program.
  
 We can get closer though to the DARPA total cost requirement by taking instead the Falcon 1's upper stage. This has a 360 kg dry mass and 3,385 kg propellant mass. The vacuum thrust is 31 kN and vacuum Isp, 330 s. Then the Calculator gives a payload of 959 kg. Taking the cost of the Falcon 1 upper stage as 1/4th that of the $8 million cost of the Falcon 1, this puts the total cost as $6.5 million

 This is a little below the DARPA requirement to LEO of at least 3,000 lbs and at a cost a bit above the $5 million limit, but likely tweaking the sizes of the lower and upper stages can get them within the required range.

 In regards to changing the size, an ideal solution would be to get an upper stage from a scaled down X-33. This would in fact allow us to get a fully reusable two-stage system. Say we scaled down the size of the X-33 by a half in the linear dimensions. This would give us a vehicle 1/8th as large in mass. Then the dry mass would be 4,000 kg with 12,000 kg propellant mass. Take the thrust as 1/8th as large as well at 300 kN, while using the same Isp 436.5 s. Then the Calculator gives us a payload of 1,902 kg.

 Given its 1/8th as large mass, we may estimate the cost to build this half-scale X-33 as $45 million. Using again a 25% price markup over 100 flights, that would be $560,000 per flight. This then would be quite close to the total cost range requirement for the DARPA program.


   Bob Clark

Saturday, September 28, 2013

Free your mind, and the rest will follow.

Copyright 2013 Robert Clark 

The story has been told that when the Native Americans first saw the ships of the Europeans they could not grasp what they were seeing because it was so outside their experience. I've always been dubious of that story. But a recent study suggests something of this nature can happen:


Science confirms: Politics wrecks your ability to do math.

By Chris Mooney
Everybody knows that our political views can sometimes get in the way of thinking clearly. But perhaps we don’t realize how bad the problem actually is. According to a new psychology paper, our political passions can even undermine our very basic reasoning skills. More specifically, the study finds that people who are otherwise very good at math may totally flunk a problem that they would otherwise probably be able to solve, simply because giving the right answer goes against their political beliefs.
http://grist.org/politics/science-confirms-politics-wrecks-your-ability-to-do-math/

 So preconceived notions can affect your ability to reason effectively, even among the smartest among us. I'm reminded also of a brain puzzler stated on the "All in the Family" TV show during the '70s. Gloria presented to the family the following:


 A father driving his young son were in an accident and the father was killed, while the son was injured but survived. When the child was brought to the hospital, the surgeon said, "I can't operate on this boy. He's my son."


 That was a puzzler the rest of the family on the show couldn't solve then and neither could I when I first saw the episode back in the '70s. The answer of course is that the surgeon was the boy's mother. 

 With the advance of women in medicine now with most med school graduates being women that probably would not be such a great puzzle to solve now as then. But it indicates how your preconceived ideas can limit your ability to solve really simple problems.
  
 Something like this is currently occurring at NASA. The Constellation program that would have returned us to the Moon has been cancelled due to high cost. However, many space advocates in the public and in Congress would prefer us to return to the Moon rather than the asteroid mission NASA is embarking on. No doubt because of these calls to return to the Moon, NASA released a study on a return to the Moon without Constellation:

Dual SLS launch campaign required for NASA’s Lunar return.

August 21, 2013 by Chris Bergin
http://www.nasaspaceflight.com/2013/08/dual-sls-required-nasas-lunar-landing-option/

 I was surprised to read that the study assumed an Altair-sized lander at the ca. 45 mT range. But the Altair's size was a big reason driving Constellation's large size and therefore great expense. And in fact by using two SLS launches the mission size in this study turns out to be even larger than Constellation. 


 It was as if the study authors had never heard of the Apollo lander that was only one-third the size of Altair. The misperception that a lunar lander has to be as large as the Altair as well as being built from scratch rather than using existing propulsive stages and crew capsules drives the false conclusion that an additional $10 billion expense would be needed for such a lander, and therefore a lunar return is unaffordable


  A further misperception is what is the mass that could be transported to LEO by the SLS. The Block 0 version of the SLS was supposed to use three SSME's on the core and use the standard 4-segment SRB's used on the shuttle. This would have a 70 mT payload capacity to LEO.


 However, NASA decided to bypass the Block 0 and go directly to the Block 1. This would stretch the core tank by a third and use a fourth SSME. It would also use a fifth-segment on the SRB's. So the size and thrust of the core would be increased by 33% and the size and thrust of the SRB by 25%.


 Despite these increases in both size and thrust, NASA was still quoting 70 mT capacity for the Block 1 SLS. Logically the payload should have been increased but NASA continued to quote 70 mT. Finally, NASA did release a report that acknowledged the payload to LEO would be 90+ mT:


SLS Dual Use Upper Stage (DUUS).

http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20130013953_2013013757.pdf


  This is important because at 90+ mT it is much easier to do a manned lunar landing mission using a single launch of the SLS, assuming you use a lander at the Apollo scale not the Altair scale. Indeed it would be possible at the first launch of the SLS in 2017.


 Then it was these preconceived notions that prevented NASA from seeing that we can in fact return to the Moon as early as 2017, and not even at significantly greater expense than that already being spent on the SLS and Orion capsule.


  Another mental block is operating in regards to how much such BEO missions should cost. NASA's commercial space program has been a great success in producing both launchers and spacecraft at as much as a 90%(!) savings over what NASA would normally have to pay for them. If any other federal agency had managed to reduce costs for normally multi-billion dollar programs to only a few hundred million dollars this would be hailed to the skies as a remarkable success in reducing costs to the American tax payer. Yet NASA was regarding it as if it were something they were only allowed to talk about in hushed tones.

 Finally, NASA has released a report detailing the savings possible under the commercial space approach:

The Commercial Leverage Model and Public/Private Partnerships.
Daniel J. Rasky
Director, Emerging Commercial Space Office
NASA Ames Research Center
Founder & Director, Space Portal
NASA Research Park
Moffett Field, CA 9403
September 11, 2013
https://dl.dropboxusercontent.com/u/47645641/AIAA_2013.pptx 

 Imagine then these cost savings applied also to BEO missions to the Moon or asteroids. This would make these missions much more fiscally feasible. It was NASA not officially acknowledging such cost savings that made it so that they could not study possibilities for returning to the Moon in a low cost fashion.

 For return to the Moon missions conducted by NASA, NASA may initially choose to use the, still expensive, SLS launcher. However, just as NASA has realized commercial space can make flights to the ISS much more cheaply than the shuttle, so also can commercial space make flights to the Moon much more cheaply.

 Indeed, by going small, going commercial, and using preexisting propulsive stages and crew modules, crewed and cargo flights to the Moon can be made for comparable costs to what we are paying the Russians to send a crew of three to the ISS.  

 The conclusion you draw is that a Moon base can be sustained on the Moon for what we are currently paying to sustain the ISS.

 Just free your mind, and the rest will follow.


   Bob Clark

Budget Moon Flights: Ariane 5 as SLS upper stage, page 2.

Copyright 2013 Robert Clark

 In the blog post Budget Moon Flights: Ariane 5 as SLS upper stage, I noted that using an Ariane 5 core stage would be a quick and low cost means of getting a higher payload capacity than the 70 mT NASA was giving for the 2017 SLS first launch. This would make it much easier to conduct a manned lunar lander mission using the SLS. However, finally, NASA has acknowledged that the Block 1, first version of the SLS to launch in 2017 will have a 90+ mT payload capacity not the 70 mT always stated by NASA:

SLS Dual Use Upper Stage (DUUS).
http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20130013953_2013013757.pdf

 This is important since it means we will have the capability to do manned lunar landing missions by the 2017 first launch of the SLS:

SLS for Return to the Moon by the 50th Anniversary of Apollo 11, page 5: A 90+ metric ton first launch of the SLS.
http://exoscientist.blogspot.com/2013/09/sls-for-return-to-moon-by-50th.html

  Still, because of the even higher payload possible using an upper stage, NASA might still want to consider getting a quick and low cost upper stage by using the Ariane 5 core. In that "SLS Dual Use Upper Stage (DUUS)" report are given some specifications for the DUUS. They state they want a highly weight optimized stage, which the Ariane 5 core would certainly fulfill. The Vulcain engine though has thrust size about twice that of the specified 100 - 120K lb range, and the 170 mT of propellant is also larger than the specified 130 mT propellant load. However, these are so much superior to the specified requirements it should result in significantly greater payload delivery than the stated 130 mT to LEO, perhaps to the 150 mT range.
 However, a problem is the 18.3 m specified max height. The Ariane 5 core is at 30.5 m height. Likely the height limitation is coming from limitations on the size and height of the facilities during stage integration. I'll find out if that is a firm limit.
 We could cut down the size of the Ariane 5 core to make it 130 mT in propellant load. Proportionally this would bring the height down to about 20 m, closer to the max. height. It turns out that shortening a stage is rather easy technically so this should still be doable by the 2017 first launch of the SLS.
 Another possibility would be to use the same propellant tank tooling for this DUUS stage as that used by NASA to make the 8 meter wide SLS tank, while using the other components such as the engine of the Ariane 5. However, the idea is to get a low cost upper stage in a short time frame. This might be a costly modification that might also be difficult to manage by the 2017 SLS first launch.
 This possibly though could open up an additional means of lightweighting the stage. We could use aluminum-lithium alloy instead of the standard aluminum used for the Ariane 5. NASA is also planning to use standard aluminum for the SLS core. But as I discuss in SLS for Return to the Moon by the 50th Anniversary of Apollo 11, page 4: further on lightweighting the SLS core, new versions of Al-Li have been developed that could be used for this purpose.
 However, due to the natural inertia of large agencies it might be difficult to change the decision and go with Al-Li for the SLS core. But this decision might be easier to make in regards to the smaller upper stage which wouldn't need as much of the more expensive Al-Li and for which getting a lightweight stage is a much greater priority.

  Bob Clark


Sunday, September 22, 2013

SLS for Return to the Moon by the 50th Anniversary of Apollo 11, page 5: A 90+ metric ton first launch of the SLS.

Copyright 2013 Robert Clark

 Finally someone at NASA acknowledges that the Block 1, first version of the SLS to launch in 2017 will have a 90+ mT payload capacity not the 70 mT always cited by NASA:

SLS Dual Use Upper Stage (DUUS).
http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20130013953_2013013757.pdf

 This is important since it means we will have the capability to do manned lunar landing missions by the 2017 first launch of the SLS.
 As discussed in the blog post SLS for Return to the Moon by the 50th Anniversary of Apollo 11, page 2: Orion + SEV design, a 90+ metric ton launcher means we can even use the Orion as the crew capsule. This is important for political reasons since the great expense spent on it means there would be a great desire among its supporters to see it be used. 
 There also is a preference at NASA for the departure stages from the lunar surface and from lunar orbit to use hypergolics, which have the surety of igniting on contact. Then another advantage of a 90+ mT SLS is that the heavier hypergolics can be used for these stages rather than the lightweight hydrogen-fueled stages I suggested in that blog post. In an upcoming post I'll show using existing hypergolic stages how we can get a lunar landing mission at less than 90 mT to LEO.
  For any of these methods it is important to use currently existing stages rather than developing them from scratch. A big reason that NASA ruled out a return to the Moon was because of the assumption that it required the development of new Altair-sized lander at a $10 billion development cost. But the need for a 45 mT Altair-sized lander is provably false as shown by the Apollo lander at one-third the size. And simply adapting already existing stages reduces the cost to a fraction of that needed for an Altair.
 So NASA is making expensive policy decisions such that we can't return to the Moon based on provably wrong assumptions. One is that the Block 1 would only have a 70 mT payload capability and so would require an expensive upper stage to increase the payload to do lunar missions, and another is that a lunar lander would require an additional $10 billion development.
 In fact, once you recognize the, obvious, fact that a lunar mission does not require an Altair-sized lander then so many possibilities become apparent. We did not have the great variety of existing launchers back in the Apollo days that we have now. If you allow your lander to be at or smaller than the Apollo lander then there are a variety of launchers that could be used for lunar missions, not just the SLS. And since they are already existing, or will be soon such as the Falcon Heavy, there would be no huge, multi-billion dollar development cost to use them. 
 So likewise also is the case for the in-space stages needed. They are already currently existing and would require relatively minor adaptations to be used for a lunar lander, for example.
 Indeed we could do manned lunar missions for what NASA is currently paying the Russians to send a crew of 3 to the ISS. The implications of that are jarring: we could have regular manned flights to the Moon for the same amount as what we are currently paying to send regular manned flights to the ISS.  And since the cargo flights to the Moon would be similarly low cost and using Bigelow style lightweight habs would allow a habitation module to be sent to the Moon on a single flight, we could have a manned lunar base for the same amount as what we are paying to sustain the ISS.
 All this comes from simply the mental reset that a lunar mission does not require the $10 billion Altair.
 Free your mind, the rest will follow.

    Bob Clark

Note: thanks to M. Moleman for discussing the NASA report "SLS Dual Use Upper Stage (DUUS)" on his blog.

Saturday, September 14, 2013

Budget Moon flights: will Canada and Europe take us back to the Moon?


Copyright 2013 Robert Clark




Canada and Europe want to send manned missions to the Moon, despite NASA's disinterest:

Canadian on Moon possible under latest space plan.
Roadmap for future missions includes lunar space station and trips to Mars.
The Canadian Press
Posted: Aug 25, 2013 8:39 AM ET
http://www.cbc.ca/news/canada/story/2013/08/25/tech-space-station-canadian-on-moon.html

In the blog post Medium Lift Circumlunar Flights, I noted that current medium lift rockets such as the Atlas V 401, Delta IV Medium, Falcon 9 could launch a Cygnus-sized capsule on a circumlunar flight. Then if the Cygnus were provided with a heat shield and life support this would provide a low cost means of performing a manned circumlunar flight.
 Orbital Sciences is investigating giving the Cygnus a heat shield based on the inflatable ones NASA is developing:


>


  Another possibility for the heat shield would be to give the Cygnus capsule the same degree of small taper as the Dragon capsule. Then you could use the same type of PICA material, which was invented by NASA Ames, as used on the Dragon.
 Still another possibility is suggested by what SpaceX is proposing for the reentry of the upper stage of a reusable Falcon 9. The stage is a cylindrical structure, but according to the images released by SpaceX, heat shield material, presumably their PICA-X material, is applied to the entry end of the upper stage and partially along one side.



 Another possibility would be to use the capsule originally designed by Andy Elson for SpaceX, called "Magic Dragon". This was to be carried by the smaller Falcon 5 and only carry 3 crew. Since the Falcon 5 had half the payload capability to the Falcon 9 and the crew size was half of the current Dragon, this smaller version likely was also half-size, at ca. 2 mT.



 In the blog post Budget Moon Flights I discussed that two cryogenic in-space stages about half-size to the Centaur could take a Cygnus-sized capsule to an actual lunar landing and back. It will actually even be possible to do it with just one of these stages: the Falcon Heavy, according to Elon Musk, will be able to send 35,000 lb, about 16,000 kg to TLI.
 Use the same Ariane 4 H10-3 upper stage used in the "Budget Moon Flights" post. This had a propellant mass of 11.86 mT and dry mass of 1.24 mT. As discussed there, take the round-trip delta-v of 8,650 m/s. The delta-v to TLI is about 3,150 m/s. Then 5,500 m/s would have to be supplied by this single stage. The stage could carry 2.9 mT payload to greater than that delta-v:

445*9.81ln(1 + 11.86/(1.24 +2.9)) = 5,900 m/s.

 Actually, it could take 3.4 mT to a delta-v of 5,500 m/s but we are limited to how much total mass the Falcon Heavy can take to TLI.
 According to the Astronautix page on the Ariane H10-3 the cost of the stage was only $12 million. Then this could serve for a low cost demonstration mission for the Canadian Space Agency (CSA) and ESA to launch as early as the 2014 expected first test flight of the Falcon Heavy.
 It is important that such low cost missions be done to break the mindset that any manned flights to the Moon have to involve super heavy lift rockets such as the Saturn V, Ares V, or SLS.


   Bob Clark





Monday, August 26, 2013

The Coming SSTO's: Page 2.

Copyright 2013 Robert Clark

 In the blog post, The Coming SSTO's I calculated some delta v's that suggested we already have the capability to do SSTO's with significant payload. However, here I'll provide some more accurate estimates by using Dr. John Schilling's Launch Vehicle Performance Calculator page. I'll go back to the Atlas rocket SLV-3 Atlas / Agena B. The specifications are given here:

SLV-3 Atlas / Agena B.
Family: Atlas. Country: USA. Status: Hardware. Department of
Defence Designation: SLV-3.
Standardized Atlas booster with Agena B upper stage.
Specifications
Payload: 600 kg. to a: 19,500 x 103,000 km orbit at 77.5 deg
inclination trajectory.
Stage Number: 0. 1 x Atlas MA-3 Gross Mass: 3,174 kg. Empty Mass:
3,174 kg. Thrust (vac): 167,740 kgf. Isp: 290 sec. Burn time: 120 sec.
Isp(sl): 256 sec. Diameter: 4.9 m. Span: 4.9 m. Length: 0.0 m.
Propellants: Lox/Kerosene No Engines: 2. LR-89-5
Stage Number: 1. 1 x Atlas Agena SLV-3 Gross Mass: 117,026 kg.
Empty Mass: 2,326 kg.
 Thrust (vac): 39,400 kgf. Isp: 316 sec. Burn
time: 265 sec. Isp(sl): 220 sec. Diameter: 3.1 m. Span: 4.9 m. Length:
20.7 m. Propellants: Lox/Kerosene No Engines: 1. LR-105-5
Stage Number: 2. 1 x Agena B Gross Mass: 7,167 kg. Empty Mass: 867
kg. Thrust (vac): 7,257 kgf. Isp: 285 sec. Burn time: 240 sec. Isp(sl): 0
sec. Diameter: 1.5 m. Span: 1.5 m. Length: 7.1 m. Propellants: Nitric
acid/UDMH No Engines: 1. Bell 8081

http://www.friends-partners.org/partners/mwade/lvs/slvgenab.htm

 We see stage 1 called the sustainer stage has nearly a 50 to 1 mass ratio. However, the Atlas had an unusual "stage and a half" structure where engines needed to lift off from the pad were jettisoned later on in the flight, leaving only a smaller, lower thrust engine behind. This engine which is the one used in stage 1, did not have enough thrust to lift off from the pad. So as in The Coming SSTO's post,  I'll replace it with the NK-33 engine which has now flown successfully on the Orbital Sciences Antares. 
 The propellant load remains 114,700 kg as in the original Atlas but the dry mass increases to 3,086 because of the heavier engine. The vacuum Isp is 331 s for the NK-33, and the vacuum thrust is 1,638 kN. Now input these numbers into Schilling's calculator. Select "No" for the "Restartable Upper Stage?" option and Cape Canaveral for the launch site. For the orbital inclination choose 28.5 degrees to match the latitude of Cape Canaveral. Then the Calculator gives these results:

====================================================
Mission Performance:
Launch Vehicle:   User-Defined Launch Vehicle
Launch Site:   Cape Canaveral / KSC
Destination Orbit:  185 x 185 km, 28 deg
Estimated Payload:   4113 kg
95% Confidence Interval: 2860 - 5625 kg

====================================================

 This value of 4,113 kg is remarkable in being close to that of the payload capability of the full Antares at 5,000 kg, a rocket of twice the gross mass, using two stages and two of the NK-33 engines on the first stage.
 Based on this, this SSTO version could be significantly cheaper than the current Antares. Plus in being only liquid fueled, it could be used as a manned launcher. Note that Orbital already has the Cygnus capsule which with the addition of a heat shield and life support could be a manned capsule.

 The mass ratio of 50 to 1 for the original Atlas is so high it would be interesting to calculate the payload capacity if we used instead the lower Isp Merlin 1D engine. By the SpaceX page, nine Merlin 1D's have total vacuum thrust of 6,672 kN. So one is 741 kN. We will need two to lift off, at 1,482 kN vacuum thrust. The two Merlin 1D's together weigh about 330 kg less than the NK-33 case, so subtract that much from the dry mass of the NK-33 case. However, the Isp is also reduced to 311 s Isp for the Merlin:
 Then Schilling's calculator gives:

====================================================

Mission Performance:
Launch Vehicle:   User-Defined Launch Vehicle
Launch Site:   Cape Canaveral / KSC
Destination Orbit:  185 x 185 km, 28 deg
Estimated Payload:   3025 kg
95% Confidence Interval: 1952 - 4331 kg
====================================================

 It is the quite high mass ratio that leads to these rather high payload capabilities.

 SpaceX might not be inclined to support such an experiment, as they are deeply invested in keeping the Falcon 9 first stage and Merlin 1D engines. However, Orbital Sciences farms out its construction of the Antares first stage to a company in the Ukraine.  So they may be inclined to try a new stage that would at the same time prove to be a revolutionary step of creating an operational SSTO.                                                                                            

   Bob Clark