New Shepard as a booster for an orbital launcher.

Copyright 2016 Robert Clark

 Blue Origin scored another first by successfully relaunching their vertical landing New Shepard suborbital rocket:

 In the blog post "Triple Cored New Shepard as an orbital vehicle", I suggested using three cores of the New Shepard rocket with a small upper stage could form an orbital launcher. However Jonathan Goff on his blog page SelenianBoondocks raised the possibility a single New Shepard could serve as the first stage booster of an orbital rocket:

Random Thoughts: New Shepard for Pop-Up TSTO NanoSat Launch.
http://selenianboondocks.com/2016/01/random-thoughts-new-shepard-for-pop-up-tsto-nanosat-launch/

  I think it should be doable using a similar small cryogenic upper stage as for the triple-cored case. The stage I suggested there was the cryogenic upper stage of the Ariane 4, the Ariane H10-3, or one developed by Blue Origin similar to it. It had a dry mass of 1,240 kg and a propellant mass of 11,860 kg. The Isp was 445 s with a vacuum thrust of 64.8 kN. However, simply using a nozzle extension as on the RL-10B-2 can give it likewise an Isp of 462 s and vacuum thrust of 110 kN. So we'll use these values.

 To make the estimate of the payload we need the vacuum values for the Isp and thrust of the BE-3 engine. In the "Triple Cored New Shepard as an orbital vehicle" blog post I estimated these to be 360 s and 568.8 kN respectively.

 However, to loft the vehicle with the additional weight of the upper stage we'll need to increase the BE-3 thrust slightly. This should doable. For instance the SSME’s could operate at 109% of their originally rated thrust, and the Merlin 1D had a 15% thrust upgrade. So say the BE-3 vacuum thrust is increased 9% to 620 kN, keeping the same Isp.

 Now use Dr. John Schilling's payload estimator program. For the "Restartable upper stage" option check "No", otherwise the payload will be reduced. Select Cape Canaveral as the launch site and enter 28.5 for the launch inclination in 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:	1690 kg
95% Confidence Interval:	1298 - 2153 kg

"Payload" refers to complete payload system weight, including any necessary payload attachment fittings or multiple payload adapters

This is an estimate based on the best publicly-available engineering and performance data, and should not be used for detailed mission planning. Operational constraints may reduce performance or preclude this mission.

 Altitude Compensation to Increase Payload.
  As I discussed in the "Triple Cored New Shepard as an orbital vehicle" blog post, altitude compensation provides a simple, low cost method of improving payload.  For instance by attaching a nozzle extension the vacuum Isp of the BE-3 can be increased to the 462 s range of the RL-10B-2 engine. The vacuum thrust will then be increased proportionally to (462/360)*620 = 796 kN.

 Then the Schilling 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:	2324 kg
95% Confidence Interval:	1841 - 2895 kg

"Payload" refers to complete payload system weight, including any necessary payload attachment fittings or multiple payload adapters

This is an estimate based on the best publicly-available engineering and performance data, and should not be used for detailed mission planning. Operational constraints may reduce performance or preclude this mission.



  Bob Clark

UPDATE, Feb. 28, 2016:

 This considered an Ariane hydrolox upper stage as the upper stage for this New Shepard launcher. This would be problematical since it would be a direct competitor to Arianespace's Vega rocket at a much lower cost than the Vega's $35 million.

 Blue Origin very likely could develop a hydrolox upper stage that would be cheaper than the Ariane one. But that would take time and significant development cost. Instead of that, Blue Origin could produce a New Shepard derived launcher for cubesats at minimal extra development cost since the required small upper stages already exist.

 Existing upper stages that could work would be the large Star solid rocket upper stages such as the Star 63F:

Star 63F:
http://www.astronautix.com/engines/star63f.htm

 Using this for the upper stage, Schillings launch performance 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:	293 kg
95% Confidence Interval:	174 - 443 kg

 This is in the range being considered for the cubesat launchers that NASA has already awarded million dollar contracts to:

Firefly, Rocket Lab and Virgin Galactic Win CubeSat Launch Contracts from NASA.
By Caleb Henry | October 15, 2015 | Feature, Government, Launch, North America, Regional, Satellite TODAY News Feed
http://www.satellitetoday.com/launch/2015/10/15/firefly-rocket-lab-and-virgin-galactic-win-cubesat-launch-contracts-from-nasa/

 Considering the quoted prices there, this New Shepard based launcher very likely could beat these prices especially using the reusable New Shepard.

 And since the upper stage already exists, it very likely would also beat to launch these other systems still in development.

 About the quick route to operational status of this orbital rocket, I think it is significant that Blue Origin was able to beat SpaceX on a relaunch of its returned booster. The argument has been made that New Shepard is not an orbital launcher. But if Blue Origin developed this orbital launcher from New Shepard then they would be able to beat SpaceX at reusing a booster for a true orbital launcher as well.

 My opinion is SpaceX will have difficulty with getting their booster to land in reliable fashion as long as it does not have hovering ability. And because the New Shepard does have hovering ability it will be more reliable as a reusable booster.

 BTW, as Blue Origin develops its large high performance dense propellant engines, it will have the same problem as SpaceX it getting its booster to be able to hover, resulting in the same problem of reduced reliability on landing. For this reason I think Blue Origin should investigate methods of giving its large planned boosters hovering ability such as discussed here:

Hovering capability for the reusable Falcon 9, page 3: hovering ability can increase the payload of a RLV.
http://exoscientist.blogspot.com/2015/12/hovering-capability-for-reusable-falcon.html

 Surprisingly, it turns out that hovering ability when properly implemented can actually improve the the payload for a reusable rocket.

Hovering capability for the reusable Falcon 9, page 3: hovering ability can increase the payload of a RLV.

Copyright 2015 Robert Clark

 Blue Origin successfully landed their New Shepard rocket after reaching suborbital space:

 Observing the last portion of the video showing the landing, deviations from the vertical are visible but the ability to hover allowed it sufficient time to correct.

 Comparing this to the SpaceX Falcon 9 failed attempts at landing it is apparent the inability to hover for the F9 did not allow it sufficient time to make the needed corrections.

 SpaceX has said they want their next test landing to be on land at the launch site. My opinion, they might succeed on the next test or two but they will always have failures without hovering ability.

Merlins in a pressure-fed mode.
 Achieving hovering is not even difficult. In the blog post "Hovering capability for the reusable Falcon 9, page 2: Merlin engines in a pressure-fed mode?" I suggested giving the Merlin the ability to run in a pressure-fed mode. The question was whether this was technically feasible. I found in fact that this process of giving a turbopump powered engine a pressure-fed mode, called an idle mode, had been successfully tested during the Apollo days on the J-2 upper stage engine.

 In giving the J-2 an idle mode though, it was changed from the gas generator cycle that is used by the Merlin 1D to a tap-off cycle:

Rocketdyne J-2.
https://en.wikipedia.org/wiki/Rocketdyne_J-2#J-2S

 However, there is an engine that uses the gas generator cycle and has an idle mode, the LE-5 upper stage engine of the Japanese space agency:

Development of the LE-X engine.
https://www.mhi-global.com/company/technology/review/abstracte-48-4-36.html

 In this idle mode though the thrust is significantly less than at full thrust, only 3% in the LE-5 case. If it is a similar low percentage for the Merlin's then all 9 engines would have to be used in this idle mode to allow it to hover on landing.

 The idle mode has an additional advantage since it does not use the turbopumps. It could be used to burn both residual liquid propellant and gases in the tanks. This would mean much less residual fluid would be left in the tank. This then reduces the amount of propellant that needs to be kept on reserve for the landing.

 Elon Musk has also recently said in his Twitter account that the F9 first stage has single-stage-to-orbit (SSTO) capability. For an SSTO the residuals in a first stage can subtract a significant amount from the payload it can deliver to orbit. Then the ability to run in an idle mode with minimal residuals left over can significantly increase the payload for an SSTO. So this would be a further advantage of giving the Merlins an idle mode.

Hovering by use of flexible nozzle extensions.
 In the blog post "Altitude compensation attachments for standard rocket engines, and applications", I discussed another method of achieving hovering capability, attaching nozzle extensions to the bottom of the engines that would allow restriction of the thrust. The flexible high temperature materials already exist in the reentry materials used in NASA's Inflatable Re-entry Vehicle Experiment (IRVE). This has the advantage that the nozzle extension would have to only be applied to the one central engine to reduce its thrust on landing.

 However, the extendable nozzle attachments also have an advantage to the SSTO case. By using an extension that can be retracted at launch and fully extended at high altitude, you can get engines usable at sea level that can reach the high vacuum Isp's usually reserved for upper stage engines. In this way the 311 s vacuum Isp of the Merlin 1D can be raised to the same level of 340 s as the Merlin Vacuum. An increase in the vacuum Isp to this extent can as much as double the payload of a SSTO.

 Note that both of these techniques, idle mode or flexible nozzle extensions, would mean hovering capability can actually increase the payload rather than reduce it.

    Bob Clark

Merlin 1A engine for a hovering Falcon 9 v1.1 first stage.

Copyright 2014 Robert Clark

  Because of the high thrust of the Merlin 1D engine and the lightweight of the reusable Falcon 9 v1.1 first stage, the stage can not hover on its return to Earth. The firing of the engine has to be precisely timed so that the rocket reaches about zero relative velocity to the Earth once it reaches the landing point. SpaceX has referred to this non-hovering mode of landing as "hover-slam". It is due to the fact the thrust to weight ratio of the stage is still above 1 with single Merlin 1 D firing when the stage is nearly empty on landing:

More on Grasshopper’s “Johnny Cash hover slam” test.
http://www.newspacejournal.com/2013/03/09/more-on-grasshoppers-johnny-cash-hover-slam-test/

 Still for safety reasons I would prefer a stage that did have the capability to hover. One concern without the ability to hover for example would be unexpected large changes in wind speed and direction known as wind shear. Airline pilots know these when they are low to the ground as "microbursts". 

Microburst schematic from NASA. Note the downward motion of the air until it hits ground level, then spreads outward in all directions. The wind regime in a microburst is completely opposite to a tornado.

 These can be potentially dangerous for pilots during takeoffs and landings since they result quickly in a large change in the aircraft's apparent airspeed, important for maintaining lift. There have been several airline accidents with wind shear identified as the cause. 

While wind shear is particularly dangerous for aircraft when near to the ground because it gives the pilots limited time to react, it does also occur at altitude. For both Space Shuttle accidents wind shear is suspected to have been a contributing factor. For the Challenger accident wind shear occurred about the same time as the shuttle reached Max Q, maximum aerodynamic pressure. This may have increased the stresses on the vehicle leading to a breach in the solid rocket boosters. In the case of Columbia, unusually strong wind shear occurred also close to Max Q that might have weakened the wing before the impact of the insulating foam.

  Recently, SpaceX had to destroy its Falcon 9R test vehicle during its last test flight:

  SpaceX has not released the cause of the accident but the rocket appeared to pitch over during the flight. There could be variety of reasons for this and not wind, but unexpected wind changes could cause it.

 The ability to hover gives you more leeway about where you land and some leeway when. You could then avoid the wind shear like an aircraft doing a go-around.

 So how to give it the ability to hover? One way would be to use a smaller engine for the landing engine. In fact SpaceX already has it in its inventory: the original Merlin 1A.

 The page on the Falcon 1 by Ed Kyle gives the Merlin 1A engine a sea level thrust of 34,900 kgf (kilograms-force). And Kyle's page on the Falcon 9 v1.1 gives the total sea level thrust using 9 Merlin 1D engines as 600,000 kgf. So one would be 66,000 kgf. Then replacing the Merlin 1D with the 1A would result in a loss of 31,000 kgf thrust. This is only a 5% loss of sea level thrust.

 Kyle's page on the F9 v1.1 though gives it dry mass of 19 metric tons (mT). Typically rocket engines leave some residual propellant left in the tanks at about 0.5%. This would be about 2,000 kg.This would give the first stage a mass at landing at about 21 mT. Then the Merlin 1A would need to be throttleable down to 60%.

 However, the Merlin 1D was made to be throttleable down to 70%, but the Merlin 1A never was. Then for this method to work the Merlin 1A would also need to be made throttleable.


       Bob Clark

Update, October 13, 2014:

 A correspondent to my Facebook page named David Whitfield suggested the possibility exhaust from the preburner alone for the Merlin 1D might be low enough to give the Falcon 9 first stage hovering capability. You might be able to use 1 to all 9 Merlin 1D preburners to provide the needed thrust.

 BTW, is this the same as the turbine exhaust that appears on the left on this image:

 UPDATE, October 25, 2014:

 Another suggestion for achieving low throttleability is to use variable size nozzles. This is discussed here:

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

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

http://www.flightglobal.com/news/articles/adventure-star-45070/

 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.

http://books.google.com/books?id=DUkl5bH6k6EC&pg=PA106&#v=onepage&q&f=false

 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

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

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

Copyright 2012 Robert Clark 

Credit: Ed Kyle's Falcon 9 page.

  I've been arguing that SSTO's are actually easy because how to achieve 
them is perfectly obvious: use the most weight optimized stages and 
most Isp efficient engines at the same time, i.e., optimize both 
components of the rocket equation. But I've recently found it's even 
easier than that! It turns out you don't even need the engines to be 
of particularly high efficiency. 
SpaceX is moving rapidly towards testing its Grasshopper scaled-down 
version of a reusable Falcon 9 first stage: 

Reusable rocket prototype almost ready for first liftoff. 
BY STEPHEN CLARK 
SPACEFLIGHT NOW 
Posted: July 9, 2012 
http://www.spaceflightnow.com/news/n1207/10grasshopper/ 

SpaceX deserves kudos for achieving a highly weight optimized Falcon 9 
first stage at a 20 to 1 mass ratio. However, the Merlin 1C engine has 
an Isp no better than the engines we had in the early sixties at 304 
s, and the Merlin 1D is only slightly better on the Isp scale at 311 s. 
This is well below the highest efficiency kerosene engines (Russian) 
we have now whose Isp's are in the 330's. So I thought that closed 
the door on the Falcon 9 first stage being SSTO. 

However, I was surprised when I did the calculation that because of 
the Merlin 1D's lower weight, the Falcon 9 first stage could indeed be 
SSTO. For the calculation we'll need the F9 dry mass and propellant 
mass. I'll use the Falcon 9 specifications estimated by GW Johnson, a 
former rocket engineer, now math professor: 

WEDNESDAY, DECEMBER 14, 2011 
Reusability in Launch Rockets. 
http://exrocketman.blogspot.com/2011/12/reusability-in-launch-rockets.html

The first stage propellant load is given as 553,000 lbs, 250,000 kg, 
and the dry weight as 30,000 lbs, 13,600 kg. 

I'll actually calculate the payload for the first stage of the new version of 
the Falcon 9, version 1.1. The Falcon Heavy will use this version's first stage 
for its core stage and side boosters. SpaceX expects the Falcon 9 v1.1 
to be ready by the end of the year. 

Elon Musk has said version 1.1 will be about 50% longer: 

Q&A with SpaceX founder and chief designer Elon Musk. 
BY STEPHEN CLARK 
SPACEFLIGHT NOW 
Posted: May 18, 2012 
http://www.spaceflightnow.com/falcon9/003/120518musk/ 

I'll assume this is coming from 50% larger tanks. This puts the 
propellant load now at 375,000 kg. Interestingly SpaceX says the side 
boosters on the Falcon Heavy will have a 30 to 1 mass ratio. This 
improvement is probably coming from the fact it is using the lighter 
Merlin 1D engines, and because scaling up a rocket actually improves 
your mass ratio, and also not having to support the weight of an upper 
stage and heavy payload means it can be made lighter. 

So I'll assume for this SSTO version of the Falcon 9 v1.1 the mass 
ratio is 30 to 1, which makes the dry mass 13 mT. 

To estimate the payload I'll use the payload estimation program of 
Dr. John Schilling:

Launch Vehicle Performance Calculator. 
http://www.silverbirdastronautics.com/LVperform.html 

It actually gives a range of likely values of the payload. But I've found 
the midpoint of the range it specifies is a reasonably accurate estimate 
to the actual payload for known rockets. 

Input the vacuum values for the thrust in kilonewtons and Isp in 
seconds. The program takes into account the sea level loss. SpaceX 
gives the Merlin 1D vacuum thrust as 161,000 lbs and vacuum Isp 
as 311 s: 

FALCON 9 OVERVIEW. 
http://www.spacex.com/falcon9.php 

For the 9 Merlins this is a thrust of 9*161,000lb*4.46N/lb = 6,460 
kN. Use the default altitude of 185 km and select the Cape Canaveral 
launch site, with a 28.5 degree orbital inclination to match the 
Cape's latitude. 

Input the dry mass of 13,000 kg and propellant mass of 375,000 kg. 
The other options I selected are indicated here:



Then it gives an estimated 7,564 kg payload mass:

===================================== 
Launch Vehicle: User-Defined Launch Vehicle 
Launch Site: Cape Canaveral / KSC 
Destination Orbit: 185 x 185 km, 28 deg 
Estimated Payload: 7564 kg 
95% Confidence Interval: 3766 - 12191 kg 
=====================================

This may be enough to launch the Dragon capsule, depending on the mass 
of the Launch Abort System(LAS). 


Bob Clark

UPDATE, Sept. 26, 2013:

 See more accurate calculations using Dr. John Schillings Launch Performance Calculator here:

The Coming SSTO's: Page 2.
http://exoscientist.blogspot.com/2013/08/the-coming-sstos-page-2.html


UPDATE, August 26, 2014:

 This blog post actually used the estimated specifications for the Falcon Heavy side boosters, as this was supposed to have an even better mass efficiency than the core stage. However, Elon Musk gave a speech where he gave some values for the Falcon 9 v1.1 first stage that allow us to estimate the propellant and dry masses for the stage. Then we can calculate the payload capability of a F9 v1.1 core stage SSTO itself. I estimate it as ca. 5,000 kg:

The Coming SSTO's: Falcon 9 v1.1 first stage as SSTO, Page 2.
http://exoscientist.blogspot.com/2013/11/the-coming-sstos-falcon-9-v11-first.html