An SSTO as "God and Robert Heinlein intended".

[RGClark]

 NASA is in a quandary right now about what to do about their manned flight capability. Congress wants this reinstituted quickly but NASA says they can't do it with the money being provided by Congress.
 In this report Boeing proposes heavy lift launchers using existing components:

Heavy Lift Launch Vehicles with Existing Propulsion Systems.
Benjamin Donahue, Lee Brady, Mike Farkas, Shelley LeRoy, Neal Graham
Boeing Phantom Works,Huntsville, AL 35824
Doug Blue
Boeing Space Exploration,Huntington Beach, CA 92605
http://www.launchcomplexmodels.com/Direct/documents/AIAA-2010-2370-650.pdf

 One of the proposals is of a manned launcher with the Orion capsule using a shuttle ET propellant tank and four RS-68 engines. This does not use an upper stage but is not a single-stage-to-orbit vehicle because the final push to orbit is made by the onboard thrusters on the Orion spacecraft.
 However, it is interesting in this report comparison is made to the S-IVB upper stage on the Apollo rocket. I was reminded of a suggestion of Gary Hudson that the S-IVB would be single-stage-to-orbit with significant payload if it used the high efficiency SSME rather than the J-2 engine:

A Single-Stage-to-Orbit Thought Experiment.
Gary C Hudson
http://www.spacefuture.com/archive/a_single_stage_to_orbit_thought_experiment.shtml

 In Hudson's proposal the vehicle could lift 10,360 lbs, 4,710 kg. This would be just enough to carry the crewed version of the Dragon spacecraft without cargo. Boeing's proposal for a manned capsule the CST-100 might be launchable by this also since it is of comparable size and design to the Dragon:

Boeing space capsule could be operational by 2015.
BY STEPHEN CLARK
SPACEFLIGHT NOW
Posted: July 21, 2010
http://www.spaceflightnow.com/news/n1007/21boeing/

 NASA has shown in their crewed spacecraft versions to want to hearken back to Apollo in their use of capsules. This SSTO idea of Hudson would have the advantage of using a proven Apollo component that is already manrated. The SSME's are also already manrated rather than the RS-68 of the Boeing proposal.
 Because of its small small size compared to the shuttle ET propellant tank it would also be relatively low cost as well as, only needing one SSME engine. In fact it would even be smaller than the Falcon 9, Delta IV, and Atlas V expendable launchers. Note as well NASA is leaning now to using SSME's or their expendable versions rather than the RS-68 for their shuttle derived manned launchers.
 Hudson in his article stated the S-IVB was designed by the Douglas Aircraft Company, which merged with McDonnell Aircraft to form McDonnell Douglas. It is notable as well that McDonnell Douglas was also the contractor on the DC-X, legendary for its low development cost, quick turnaround time, and small ground crew.
 NASA in their shuttle-derived launcher studies have focused on getting a cheaper version of the SSME by making an expendable version. However, the greatest advantage of a SSTO is in being reusable. Then I suggest studies be made on the SSME going the opposite direction: how can it be made to be reusable at much reduced maintenance cost?
 Now the SSME's have to be overhauled after every flight costing ten's of millions of dollars. However, Henry Spencer a highly regarded expert on the history of space flight has said Rocketdyne studies show that with a lot of work to upgrade it, maintenance could be reduced to $750K per flight:

Engine reusability (Henry Spencer)
http://yarchive.net/space/rocket/engine_reusability.html

Spencer here said this would not be satisfactory for really large reductions in space costs. But this would be a reduction in SSME maintenance costs by 1 to 2 orders of magnitude, a major reduction in the costs for using the engine. The question is: how much would it cost to make the necessary upgrades to the engine?


   Bob Clark

[ijuin]

For an expendable rocket, the difference between true SSTO (the booster provides the orbital insertion burn) and almost-SSTO (the booster drops off just short of orbit and the payload performs the orbital insertion burn, as with how the Shuttle Orbiter performs its own insertion burn after jettisoning the External Tank) is mostly academic. In practice, it is probably preferable to drop the booster in the Pacific after a single orbit as opposed to spending the extra fuel to boost it that last 100-200 m/s to put it into orbit and fall into the atmosphere who-knows-where.

Anyway, if a single-stage design can be made as cheaply and with as much payload capacity as Falcon 9, then I am in favor of its use, especially if it is an all-hydrolox design that eliminates the toxic exhaust that solid or hypergolic fuels produce.

[RGClark]

Another option for a manned launcher. In this report Boeing proposes heavy lift launchers using existing components:
Heavy Lift Launch Vehicles with Existing Propulsion Systems.
Benjamin Donahue, Lee Brady, Mike Farkas, Shelley LeRoy, Neal Graham
Boeing Phantom Works,Huntsville, AL 35824    
Doug Blue
Boeing Space Exploration,Huntington Beach, CA 92605
http://www.launchcomplexmodels.com/Direct/documents/AIAA-2010-2370-650.pdf
 One of the proposals is of a manned launcher with the Orion capsule using a shuttle ET propellant tank and four RS-68 engines. This does not use an upper stage but is not a single-stage-to-orbit vehicle because the final push to orbit is made by the onboard thrusters on the Orion spacecraft.
 However, it is interesting in this report comparison is made to the S-IVB upper stage on the Apollo rocket. I was reminded of a suggestion of Gary Hudson that the S-IVB would be single-stage-to-orbit with significant payload if it used the high efficiency SSME rather than the J-2 engine:
A Single-Stage-to-Orbit Thought Experiment.
Gary C Hudson
http://www.spacefuture.com/archive/a_single_stage_to_orbit_thought_experiment.shtml
 In Hudson's proposal the vehicle could lift 10,360 lbs, 4,710 kg. This would be just enough to carry the crewed version of the Dragon spacecraft without cargo.

It is notable that the upper stage of the Ares I is based on this S-IVB stage. Then this upper stage as well should be able to act as an SSTO with an SSME engine. This is important because the Ares I upper stage was originally planned to use the SSME, so this means much of the technical and financial analysis of using the SSME for the upper stage of the Ares I has already been done.
 However, because of the cost of the SSME engine and technical risk in making it airstartable, the decision was made to use the J-2X engine instead. But for the SSTO purpose you don't have the problem of making it airstartable, and as I discussed the reusability maintenance costs can be reduced by an order of magnitude for the SSME.
 This report contains some of the specifications on the Ares I upper stage:

NASA’s Ares I Upper Stage.
http://www.nasa.gov/pdf/231430main_UpperStage_FS_final.pdf

 The propellant mass is listed as 138 mT, the dry mass of the stage as 17.5 mT, and the interstage mass, as 4.1 mT. See the second attached image below taken from page 2 of the report. The interstage supports the weight of the upper stage on top of the lower stage so won't be needed for the SSTO version. So we can take the dry mass now as 13.4 mT.
 We need to add onto this now the extra weight of using the SSME over the J-2X engine. The report lists the J-2X mass as 2.5 mT. The SSME mass is 3.1 mT, .6 mT heavier. This brings the dry weight to 14 mT.
 A puzzlingly high value of 2.5 mT however is given for the avionics. You wouldn't think it would need to be this high if it consisted of just electronics and computer systems with modern miniaturization. Most of the avionics is included in the "instrument unit". As you can see from the first attached image below, the instrument unit is regarded as a separate element of the upper stage and is contained within the forward skirt of the stage. The forward skirt serves to support the weight of the Orion CEV, so needs to have significant strength and mass to support the 20,000+ kg weight of the Orion spacecraft. So I'm wondering if that 2.5 mT mass is including the mass of this forward skirt.
 The forward skirt mass can certainly be reduced if using a Dragon spacecraft at only one quarter the mass of the Orion. So that part of the dry mass will be reduced, though it's uncertain if the avionics mass itself can be reduced. In any case using 14 mT dry mass of the Ares I upper stage, the 138 mT propellant mass, the 425 s trajectory averaged Isp of the SSME given by Hudson, and a 9,200 m/s required delta-V to orbit, we can calculate the payload to orbit can be 3 mT:

425*9.8ln(1 + 138/(14 + 3)) = 9,205 m/s.

  This payload mass would not be enough for the Dragon spacecraft but might be enough for an innovative new spacecraft proposal from the University of Maryland:

Phoenix: A Low-Cost Commercial Approach to the Crew Exploration Vehicle.
http://www.nianet.org/rascal/forum2006/presentations/1010_umd_paper.pdf

 This uses a cylindrical shape for the capsule so would have more space for the crew/passengers. It also uses a new design for a thermal protection system called a "parashield" that would save weight over the traditional ablative design. The mass of the capsule in this study is given as 3,268 kg, so would only have to be reduced by a small proportion to fit within the payload mass constraints.
 However, it might be possible to increase the payload capability of the SSME-powered Ares I upper stage to be able to carry even the Dragon spacecraft. First, more propellant can be carried in the same size tanks by densifying the propellant by subcooling:

Liquid Oxygen Propellant Densification Unit Ground Tested With a Large-Scale Flight-Weight Tank for the X-33 Reusable Launch Vehicle.
http://www.grc.nasa.gov/WWW/RT/RT2001/5000/5870tomsik.html

 As much as 10% more propellant can be carried by subcooled densification. This corresponds to 10% greater mass that can lofted to orbit. So from a 17 mT total of launch vehicle + payload, up to 18.7 mT. This extra mass can go to extra payload so to 4.7 mT payload.
 Secondly, recent research has shown that from 10% to 20% weight savings can be made off the structural weight on launch vehicles:

NASA Recalculates To Save Weight On Launchers.
Jan 5, 2011
By Frank Morring, Jr.
http://www.aviationweek.com/aw/generic/story.jsp?id=news/awst/2011/01/03/AW_01_03_2011_p53-277413.xml&headline=NASA%20Recalculates%20To%20Save%20Weight%20On%20Launchers&channel=space

The structural mass sans engine is 11 mT. If 10% weight can be saved off this then that can be transfered to extra payload, bringing the payload capacity to 5.8 mT. This would then be within the payload capacity to carry the Dragon spacecraft.


  Bob Clark



Images:
 
Ares I elements.
http://oi56.tinypic.com/5voinb.jpg

Ares I upper stage.
http://oi55.tinypic.com/1bxvr.jpg

[RGClark]

Another option for a manned launcher. In this report Boeing proposes heavy lift launchers using existing components:
Heavy Lift Launch Vehicles with Existing Propulsion Systems.
Benjamin Donahue, Lee Brady, Mike Farkas, Shelley LeRoy, Neal Graham
Boeing Phantom Works,Huntsville, AL 35824   
Doug Blue
Boeing Space Exploration,Huntington Beach, CA 92605
http://www.launchcomplexmodels.com/Direct/documents/AIAA-2010-2370-650.pdf
 One of the proposals is of a manned launcher with the Orion capsule using a shuttle ET propellant tank and four RS-68 engines. This does not use an upper stage but is not a single-stage-to-orbit vehicle because the final push to orbit is made by the onboard thrusters on the Orion spacecraft.
 However, it is interesting in this report comparison is made to the S-IVB upper stage on the Apollo rocket. I was reminded of a suggestion of Gary Hudson that the S-IVB would be single-stage-to-orbit with significant payload if it used the high efficiency SSME rather than the J-2 engine:
A Single-Stage-to-Orbit Thought Experiment.
Gary C Hudson
http://www.spacefuture.com/archive/a_single_stage_to_orbit_thought_experiment.shtml
 In Hudson's proposal the vehicle could lift 10,360 lbs, 4,710 kg. This would be just enough to carry the crewed version of the Dragon spacecraft without cargo.

The point of the matter is that if you use highly weight optimized structures and high efficiency engines at the same time then what you wind up with will be a SSTO capable stage. The Ariane 5 core stage is another weight optimized structure using common bulkhead design for its propellant tanks. The Ariane 5 core stage will also become SSTO if using high efficiency SSME's instead of the Vulcain engines.
 The specifications of the Ariane 5 are given here:

Ariane 5 Data Sheet.
http://www.spacelaunchreport.com/ariane5.html

 The Ariane 5 generic "G" version could be lofted by a single SSME. It's gross mass is listed as 170 mT, and the propellant mass as 158 mT, for a dry mass of 12 mT. The Vulcain engine is listed on this page as weighing 1,700 kg:

Vulcain - Specifications.
http://www.spaceandtech.com/spacedata/engines/vulcain_specs.shtml

  Switching to a heavier SSME engine would add 1.4 mT to the vehicle dry mass, so to 13.4 mT for the dry mass. Using a 425s average Isp again for the SSME, this would allow a 6,000 kg payload:

425*9.8ln(1 + 158/(13.4+6)) = 9,218 m/s.
 
 We wish to use this for a man-rated vehicle though. The Ariane 5 was originally intended to be man-rated using the Hermes spaceplane to carry crew. However, it's not certain the degree this was followed-through when the Hermes was canceled.
 As with the Ares I upper stage, there are means to increase the payload capacity. Subcooled densification allows 10% greater propellant to be carried, so then 10% greater mass can be lofted to orbit. This brings the total lofted weight from 19.4 mT to 21.3 mT. This extra weight can go to extra payload, so from 6 mT to about 8 mT in payload.
 The Ariane 5 uses an aluminum alloy, but not the aluminum-lithium alloy being used now for the lightest weight designs.  Switching to aluminum-lithium allows approx. 10% weight saving over the previous aluminum alloy. The structural mass sans the SSME engine is 10.3 mT, so about 1 mT would be saved that could go to extra payload.
 I also mentioned before the new research that suggests 10% to 20% can be saved in structural mass because of overly conservative design now used. This would be another 1 mT that could be saved off the dry weight. These weight savings could go to extra payload, bringing the payload capacity to 10 mT.
 ESA appears to be amenable to adapting the Ariane 5 core stage for other uses, considering its agreement with ATK to use it for an upper stage. So NASA or a private company should be able to make an agreement with the ESA to use it for this purpose, based on getting sufficient financing. In this regard, to get a prototype done at low cost I suggest using the RD-0120 russian analogue of the SSME's. These are in mothballs and probably can be obtained at greatly reduced price. As a point of comparison the NK-33 was mothballed by the russians and Aerojet was able to buy 36 of them for only $1.1 million each(!) Aerojets version of the NK-33 is now on track to be used by Orbital Sciences on their Taurus II launcher.
 Then the Ariane 5 core version of this SSTO has the advantage over the Ares I upper stage and S-IVB versions in being already built and in current use. It also has now the capability when powered by an SSME or RD-0120 to launch a SpaceX Dragon sized spacecraft to orbit without having to use special fuel densifying or lightweighting methods.
 NASA has said they want to support commercial space. Support for this launcher would allow for a small, relatively low cost launcher that would permit independent private companies to launch their own manned, or cargo flights to space.



  Bob Clark