SpaceX BFR tanker as an SSTO.

Copyright 2017 Robert Clark

 Elon Musk has suggested the development of orbital point-to-point manned transport may pay for the development of his Mars colonization plans, the idea being there would be a great market for such manned flights. Peter Diamandis has made this point as well, that there would be a great market for such flights to orbit:

Peter Diamandis: Taking the next giant leap in space.

 Elon suggests such orbital transports would be best implemented as two-stage vehicles. However, this simulation shows the upper stage of the SpaceX Interplanetary Transport System (ITS) introduced by Elon Musk in 2016 could get a total 190 metric tons to orbit as an expendable SSTO, including the stage and the payload:

 Since the stage was estimated to weigh 90 tons, this would mean 100 metric tons payload as an expendable SSTO. Then the question is how much mass would be taken up for propellant for return using vertical landing.

 However, in the latest incarnation introduced in 2017 the upper stage of the now BFR is about half as large in propellant load and with 6 engines instead of 9. Here are screen grabs from the video on the latest version:

  So we may estimate this half-size version to have half the dry mass at ca. 45 metric tons and could get approx. 50 metric tons to LEO as an expendable SSTO.

 As justification for this dry mass estimate, here's the description of the original ITS upper stage, both spaceship and tanker versions:

 And here's the description of the BFR spaceship, half size to the ITS version:

 You see the BFR spaceship is at about half the listed dry mass value as the ITS spaceship.  Actually during the video Musk says the design mass was just at half at 75 tons, but the 85 tons mass was allowing for weight growth. So it is plausible the BFR tanker is ca. half the mass of the ITS version or a little more, ca. 45+ tons.

 We'll give it 9 engines instead of 6 so it'll have enough thrust to lift off with the heavy payload, and give it a dry mass of 50 metric tons with weight growth. Now do a payload estimate using Dr. John Schilling's launch performance estimator. The high expected thrust/weight ratio for the Raptors means they'll weigh perhaps 1,000 kg each. So the addition of 3 will add perhaps 3 tons to the dry mass. Since the payload is so high this will be a relatively small payload loss.

 The Schilling estimator takes the vacuum values for the Isp and thrust, so enter 375 s as the Isp and 9*1,900 kN= 17,100 kN for the total thrust.

 Insure that the "Restartable upper stage" option is set to "No" otherwise the payload will be reduced. And set the launch inclination to match the launch site, so at 28.5 degrees for Cape Canaveral:

 Then the result is:

 Confirming the ca. 50 metric ton payload as an expendable SSTO.

 But this puts it as an expendable SSTO in the payload range of the expendable Falcon Heavy while being also in the same size range of the Falcon Heavy. So this SSTO would get the same payload fraction as a 2 and 1/2 stage vehicle. Moreover, judging from the fact the ITS tanker upper stage was to cost $130 million production cost, the half size BFR tanker might only be $65 million, so it would be half the cost of the Falcon Heavy. But the Falcon Heavy as an expendable launcher already would be a significant cut in the cost to orbit. So the BFR tanker as an expendable SSTO could be a great reduction in the cost to space, compared to current values.

 I had earlier done a calculation that showed the Falcon Heavy as an expendable with 53 metric ton payload capacity and $125 million launch cost could be financially feasible as a tourism vehicle to orbital space or transport to orbital space hotels:

Falcon Heavy for Orbital Space Tourism.
http://exoscientist.blogspot.com/2014/09/falcon-heavy-for-orbital-space-tourism.html
 ​
 So this BFR tanker could likewise be feasible financially as an expendable SSTO, as the price should be well less than $125 million as an expendable. But of course SpaceX wants to make it reusable. The reusability should cut the launch cost multiple times. Then the question is how much will reusability cut into the payload mass?

Reusable SSTO case.
  In the presentations on the ITS and BFR both the spaceship and tanker versions of the upper stage were always presented as reusable. So it is likely the heat shield mass is already included in the cited vehicle dry mass values. I'm estimating though surprisingly high values for the thermal protection of the BFR upper stages, either spaceship or tanker versions. I'm using the fact as indicated in the wiki page on the BFR that it will use the PICA-X thermal protection material. Several references give the PICA-X density as about 0.25 gm/cc = 250 kg/m3, and the thickness as on the Dragon 2 as 7.5 cm, 0.075 m, about 3 inches.

 The BFR upper stage has a length of 48 meters and a width of 9 meters. The top part of the stage is conical. Visually, this top portion is about 1/3rd the vehicle length, so about 16 meters long. So I'll approximate the bottom area to be covered by thermal protection covered as 32*9 + (1/2)*16*9= 360 m2.

 Then the volume of the thermal protection material is 360 * 0.075 = 27 m3. At a density of 250 kg/m3, that amounts to a mass for the thermal protection of 27 * 250 = 6,750 kg, which is a surprisingly high included mass in the dry mass of 85 tons for the spaceship upper stage or the 50 tons for the tanker upper stage.

 One possibility, is the thickness of the PICA-X for the Dragon 2 is coming from the fact it is doing a ballistic reentry, thus generating high heat. However, the BFR upper stage will be doing a more gentle gliding reentry. So perhaps the thermal protection will only need to be half as thick, so weigh half as much. For instance, the heat shield tiles on the underside on the space shuttle only weighed half the PICA-X tiles.

 And new versions of these space shuttle tiles used on the X-37B are more durable while remaining lightweight:

The X-37B stands in front of part of the fairing that protects it during launch, showing off the silica tiles on its underside. Those TUFI (toughened uni-piece fibrous insulation) tiles are said to be more durable than their counterparts on the space shuttle. On the leading edge of the wings, meanwhile, are TUFROC (toughened uni-piece fibrous refractory oxidation-resistant ceramic ) tiles, which NASA named the government winner of its 2011 Invention of the Year Award.

http://www.space4peace.org/articles/space_plane_of_mystery.htm

 Elon has implied the reusable version of the BFR upper stage would only get perhaps in the range of 10 to 15 metric tons payload (by saying it's an order of magnitude less than the full BFR 150 ton reusable payload.) That loss in payload is not coming from the heat shield mass since that's already included in the vehicle dry mass. The loss in payload is high though, 40 tons, nearly the size of the entire vehicle dry mass, presumably because of the amount of the propellant that needs to be kept on reserve for landing on return.

 I'd like to see a trade study of the payload of instead going with wings for a horizontal landing. See for example the discussion here:

http://yarchive.net/space/launchers/horizontal_vs_vertical_landing.html

 Wings typically take up only 10% of an aircraft dry mass. Then with carbon composites, that would be cut to less than 5% of the landed (dry) mass. Keep in mind the loss in payload with vertical, propulsive landing is nearly 100% of the vehicle dry mass. Also, going with short, stubby wings as with the X-37B, you can make the wing weight even less:

 The areal size of the wings in that case would also be less than that of bottom area of the BFR tanker, perhaps only 1/4th to 1/3rd the areal size. So the increase in heat shield mass would only be at most 1/3 that of the approx. 6,750 kg mass of the current heat shield, so perhaps an extra 2,250 kg. But actually the addition of wings gives a gentler glide slope so probably the heat shield thickness could be reduced. The result might even be the total heat shield mass would be reduced by adding wings.

 Elon has spoken of preferring vertical landing because it could be used generally on both worlds with and without atmospheres. However, to achieve his desired goal of making mankind multiplanetary, making human orbital spaceflight commonplace is an important part of that goal. A lower development and production cost BFR upper stage acting as a reusable SSTO would go a long way towards that goal. So even if the optimized orbital or point-to-point transport looks completely different than the interplanetary lander, such as requiring wings for example, it would still be important to develop it.

 Advantage of altitude-compensation.
 In the discussion after the introduction of the BFR, Elon Musk and other commenters on various space forums, engaged in alot of speculation on the optimal combination of sea level engines, vacuum engines, and some medium, intermediate area ratio engines. This is necessitated because the highest Isp engines, optimized for vacuum, can not operate reliably, and safely at sea level. But then using sea level engines or even intermediate level engines would subtract from the Isp possible.

 This illuminates once again the importance of implementing altitude compensation engines in space flight, at least for Earth launch. This would permit the high thrust needed at launch at sea level, as well as the high Isp needed at near vacuum.

 There are many ways to implement the altitude compensation and none are particularly difficult to do. Some methods are discussed here [1], [2], [3], [4], [5].


REFERENCES.

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

2.)Altitude compensation attachments for standard rocket engines, and applications, Page 2: impulse pressurization methods.
http://exoscientist.blogspot.com/2016/01/altitude-compensation-attachments-for.html

3.)Altitude compensation attachments for standard rocket engines, and applications, Page 3: stretchable metal nozzles.
http://exoscientist.blogspot.com/2016/06/altitude-compensation-attachments-for.html

4.)Altitude compensation attachments for standard rocket engines, and applications, Page 4: the double aerospike.
http://exoscientist.blogspot.com/2016/10/altitude-compensation-attachments-for.html

5.)Altitude compensation attachments for standard rocket engines, and applications, Page 5: metal foil expandable nozzles.
http://exoscientist.blogspot.com/2017/08/altitude-compensation-attachments-for.html

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

Budget Moon flights: lightweight crew capsule.

Copyright 2013 Robert Clark

 In the post Budget Moon flights, I argued that by using a capsule half-sized to the Dragon capsule at ca. 2,000 kg dry mass, that we could launch a manned lunar landing mission carried to LEO by a single Delta IV Heavy or Ariane 5 ME, with a separate man-rated launcher to carry the capsule and crew. And if using the Falcon Heavy at the stated $1,000 per pound price point it could be done at launch costs of less than $100 million.

 I'll discuss such a half-sized capsule here. This report discusses a capsule with such low dry mass:

Phoenix: A Low-Cost Commercial Approach to the Crew Exploration Vehicle

Abstract: Since the announcement of President Bush’s Vision for Space Exploration (VSE) in early 2004, the architecture of Project Constellation has been selected. The system will be centered around the Crew Exploration Vehicle (CEV), which has been dubbed by NASA administrator Michael Griffin as “Apollo on steroids”. The CEV is to be launched on a new launch vehicle, derived from existing shuttle technology. The development of this new
spacecraft and launch vehicle is a very costly proposition. An alternate approach is proposed in this study. The Phoenix is a smaller spacecraft designed specifically to be launched on the Falcon 5 vehicle under development by SpaceX. Because the SpaceX vehicle will cost only a fraction of today’s launch costs, the Phoenix is estimated to cost less than half of the price of the CEV. This reusable three person capsule utilizes an innovative re-entry concept, which allows for a cylindrical spacecraft with greater interior volume. This extremely cost-effective spacecraft is an attractive option for fulfilling VSE requirements.

 Below is page 3 from this report:

  ...
_______________________________________________________________________

_______________________________________________________________________
  ...

 Our crew capsule would not need its own propulsion system, so get the dry mass by subtracting off the masses in the "Propulsion (SPS)" section to get 2,160 kg. We will retain the RCS system though. Note the 2,160 kg dry mass actually also includes the mass for a crew of three and the mass for their space suits, as well as the food and water for a mission to and from the Moon.
 The heat shield is of the innovative "parashield" design:

This with its truss structure/drive system would amount to about 15% of the capsules dry weight. This is about the same as Apollo era heat shields. However, the lightweight PICA-X material used on the Dragon capsule can withstand lunar return velocities and would only weigh half as much. This would subtract about 160 kg from the dry mass.

 This report on a "Phoenix" capsule only envisions this crew module to carry the crew from LEO to lunar orbit and back with a separate module to be used on a lunar lander, a la the Apollo architecture. But following the Early Lunar Access architecture we could use this one single crew module for the entire flight.

 It is notable that Orbital Sciences Cygnus capsule is of similar size to this "Phoenix" crew capsule.

Artist's rendering of Cygnus spacecraft approaching the International Space Station.

CREDIT: Orbital Sciences Corporation

 The Cygnus dry mass is 1,500 kg. This includes the propulsive service module at the base. The service module is based on Orbital's Star satellite bus. According to Astronautix this has a dry mass of about 800 kg. So the capsule itself is 700 kg. Adding on the life support elements as given in the "Phoenix" capsule report, would result in a dry mass of about 2,000 kg.
  This is important because the Cygnus is built by Thales in Italy so it means Europe could make all the components for the Moon mission, including the man-rated launch rocket as described in post The Coming SSTO's: multi-Vulcain Ariane.
 The "Phoenix" capsule report estimates billions of dollars in development cost. But this is using traditional NASA costing estimates. However, SpaceX has shown that development costs can be cut by a factor of 10 by private financing both for launchers and for crew capsules. From its half-size compared to the Dragon we might estimate its development as privately financed at half of the $300 million spent developing the Dragon, so to only $150 million.


   Bob Clark

Update, May 1, 2014:

 For the passage above highlighted in red, I am now inclined to think that 1,500 kg dry mass does not include the mass of the service module, which is, the propulsion unit at the bottom. It's because of this page:

Cygnus Spacecraft Information

Pressurized Cargo Module

Cygnus standard and Cygnus enhanced use Pressurized Cargo Modules built by Thales Alenia Space, Italy. The module is based on the Multi Purpose Logistics Module that was flown on multiple Space Shuttle Missions to deliver pressurized cargo to ISS.

It is 3.07 meters in diameter and has a length of 3.66 meters in its standard configuration and 4.86 meters with one segment added in the enhanced configuration. The standard PCM has a dry mass of 1,500 Kilograms and the enhanced version weighs 1,800kg. 2,000 Kilograms of cargo can be packed into the standard PCM while the enhanced version allows 700kg more to be loaded. The cargo module has a pressurized volume of 18.9 cubic meters in the standard configuration and 27 cubic meters in the enhanced configuration. Power consumption of the PCM is less than 850 watts.

http://www.spaceflight101.com/cygnus-spacecraft-information.html

 This page also has a separate section on the service module which suggests that mass amount above was only for the pressurized module itself.