A Kerosene-Fueled X-33 as a Single Stage to Orbit Vehicle.

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Composites can use different fabrication techniquies depending upon size and application. When used for smaller non-structural applications they can be cheap. Even in some aeronautical applications they can be made cost effectively. However, larger load bearing primary structural members would need to be bonded, which generally means Autoclave (a large pressurized oven). When fabricating the X-33, one of the largest Autoclaves in the Western Hemisphere was used to bond the members of the Composite Hydrogen tank which failed. It was 22' in diameter. Building larger pieces out of composites would require newer Autoclaves, which are very expensive in such sizes.

I mean a launch vehicle can be 200' long or more. To save weight, would imply main structural members which run the entire length without fasteners. Building a 200' load bearing composite beam to aerospace tolerances is beyond current capability. People have looked at doing this for air vehicles, and considered various fabrication methods, but it has never actually been done in an airplane let alone a space launch vehicle. Other vehicle structures might be larger than the X-33 tank, and would also be potentially un-fabricatable using composites. Again, trinkets are one thing. Even select (mostly non-load bearing) aircraft parts are a second thing. Making an all-composite launch vehcile is another thing entirely.

As far as cost. $100M is the approximate cost of a main US launcher like Atlas or Delta. $5B for development costs is a very fair number, when one considers how much Shuttle cost or the cited $1.3B figure for X-33 as a suborbital demonstrator vehicle. Recurring costs for a fleet of such vehicles might run $1B per vehicle (think Shuttle) + $100M per launch (think Atlas/Delta). So again, how many flights does it have to fly to become more economical than today's disposable. If you build 4 vehicles, and they each fly 50 times (a little more than Shuttle) then total cost is $5B + 4x$1B + 200x$0.1B = $29B for 200 flights, vs 290 flights on Atlas/Delta, so same order of magnitude.

It is this math that has lead NASA to go back to staged disposable rockets, and elad the US Air Force to goto EELV in the first place. Now in 20 years technology may change and it may become more economical to do such things like comopsite launchers, just as the Saturn rockets were canned due to poor economics 30 years ago, and subsequantly EELV was brought online. On the other hand, as I mentioned earlier, in 20 years we may be close to space elevators or Bolos or some other orbit access technology that is cheaper, safer and more reliable.


As I recall the DOD decision to abandon the Shuttle was made after Challanger, many years before the EELVs, and was made for reasons of reliability rather than cost. They switched initially to the Titan IV, which turned out to be about the same cost as the Shuttle and which proved even less reliable. It was then that the EELV program was initiated, a competition between Atlas and Delta which unaccountably led to the selection of both, resulting ultimately in both programs being underutilized.

Regarding the practicality of reusable launch vehicles, the Shuttle was the first, and so far the only, reusable launch vehicle/spacecraft ever built. Its cost of operation is due primarily to 1) large portions of the structure are either expendable (ET) or must be completely disassembled and rebuilt (SRBs), 2) design elements chosen early in the program resulted in a very high ratio of maintenance man-hours to flight cycles, (i.e. the thermal protection system and the electrical wiring harness with its rather friable insulation) and 3) facilities costs imposed a very high fixed overhead (VAB, Crawlers, MLPs, LC-39). None of these characteristics are intrinsic to reusable launch vehicles. That is why the technology demonstrators were critical to a new generation of RLVs; the Saturn V notwithstanding, systems analysis is no substitute for hands-on flight experience with prototypes, and they are essential if practical systems are to be developed. . While the shuttles have actrually completed about 30 flights each, they are nowhere near their structural fatigue limits. While I have seen may statements that they are getting "old and unreliable" I have yet to see any engineering data to support this. Certainly there is no reason an RLV would be intrinsically limited to any specific number of flights, or that each flight would cost $100M.

The only thing physically required for flight is energy, and the cost of all the fuel for the Shuttle is less than 1% of the mission cost, In fact the helium alone costs more than the fuel. Obviously for a new generation of reusable launch vehicles any designer would consider alternatives to helium, but again the lack of prototypes and the philosophy of creating the final design by analysis alone reslted in a complete failure to appreciate the impacts of such design decisions on operating costs.

In summary, the proper response to the high cost of spaceflight is not retreating to a design that we know is too expensive to ever be practical, it is rather to test in flight a series of prototypes that will verify or reject critical decisions in design. This is of course nothing more than the fundamental approach that has been used successfully in aviation from its earliest days.

Astro_Robert":5lgadu5w said:
Also note that the much ballyhood DC-X used a tank liner for its tanks..

I cannot find any definitive reference on this and would be very interested if youhave details. The original DC-X used metallic tanks, the DC-XA had a carbon-fiber LH2 tank. (see http://www.astronautix.com/lvs/dcx.htm ) I cannot find any reference that clearly says whether the tank had a liner, but the principal cause of the damage on the X-33 was vacuum pumping of air into the tank laminations from outside rather than hydrogen from inside, so a liner (unless it was insulating) would not have prevented the problem. I have read that the complex multilobed shape of the X-33 tank and the choice of resin systems were factors in its failure.

Finally, I believe the DC-XA tank was fabricated in two sections and bonded together; there is no reason a large composite structure must intrinsically be made in one piece. But since large composite primary structures are essential for the next generation of airliners, even the 787, which has just made its first flight, and the White Knight II, it seems inconsistent to say they are impractical for spacecraft.


On the other hand, as I mentioned earlier, in 20 years we may be close to space elevators or Bolos or some other orbit access technology that is cheaper, safer and more reliable....

Not being an engineer with a specialization in composites what I read and understand about the subject would suggest the technology simply doesn't exist, yet. 20 years ago the space elevator was just twenty years ago and hasn't budged more then a few inches in the time line. Whats a Bolos? Never heard of it.

SSTO is actually harder to do because maximizing nozzle size to altitude would cost too much to throw away and not re-use, something the X-33 nozzle is said to do, theoretically. It's better to compromise over a smaller band. First stage to altitude with maximum speed that allows the second stage to reach orbit. The second stage engine then operates any number of times in orbit and is sent back for overhaul as cargo.


moon shuttle

I'm curious to know why I haven't heard any mention of a moon shuttle. The DC-X SSTO would only have to get itself to orbit be stripped of its aerodynamic skin be refueled and then spend the rest of it's life going from the ISS to the moon and back. What am I missing that would make this possible?


scottb50":3rqi44t5 said:
On the other hand, as I mentioned earlier, in 20 years we may be close to space elevators or Bolos or some other orbit access technology that is cheaper, safer and more reliable....

Not being an engineer with a specialization in composites what I read and understand about the subject would suggest the technology simply doesn't exist, yet. 20 years ago the space elevator was just twenty years ago and hasn't budged more then a few inches in the time line. Whats a Bolos? Never heard of it.

SSTO is actually harder to do because maximizing nozzle size to altitude would cost too much to throw away and not re-use, something the X-33 nozzle is said to do, theoretically. It's better to compromise over a smaller band. First stage to altitude with maximum speed that allows the second stage to reach orbit. The second stage engine then operates any number of times in orbit and is sent back for overhaul as cargo.

There are other options of altitude compensating nozzles than the aerospike.



There are other options of altitude compensating nozzles than the aerospike.....

It is much simpler to use a TSTO vehicle, then the Upper Stage can have an optimized nozzle.


scottb50":jbrmddc3 said:
There are other options of altitude compensating nozzles than the aerospike.....

It is much simpler to use a TSTO vehicle, then the Upper Stage can have an optimized nozzle.

I agree, but we also need to investigate these possibilities.

Lets say we had advanced technology wise to a stage where an SSTO is a realistic possibility, with comparable ISP to a TSTO setup and assuming the SSTO setup was just as reliable and cost effective as a TSTO setup would you consider it?


Lets say we had advanced technology wise to a stage where an SSTO is a realistic possibility, with comparable ISP to a TSTO setup and assuming the SSTO setup was just as reliable and cost effective as a TSTO setup would you consider it?

Absolutely, but it doesn't exist and doesn't appear to be possible without many major breakthroughs. I'm thinking of what could be up and running in a couple of years.


Agreed there would be leaps in technology required in order to achieve this in two years.

How about we setup an online open source study into building one?


annodomini2":2zlzguff said:
Agreed there would be leaps in technology required in order to achieve this in two years.

How about we setup an online open source study into building one?

Fine, if that would help. Except there is no leap in technology just improving and simplified adaptions of current technology.


Here is a little improvement of composite structures in space:
Alternate Space Capsule Concept Passes Tests
A NASA team looking into design concepts for future space capsules has successfully demonstrated that an all-composite structure is a feasible alternative to traditional metal capsules for carrying astronauts into space and returning them safely to Earth.

"Our tests showed that a composite module can 'achieve the mission' with damage that is likely to occur but could go undetected," said Mike Kirsch, manager of the Composite Crew Module (CCM) project. "The test article withstood twice the design internal pressures with known damage and then was subjected to cyclic testing to four times the design life with no detrimental damage growth," he added.

Partners include subject matter experts from nine of ten NASA centers, the Air Force Research Laboratories, Alcore Corporation, Alliant Techsystems (ATK), Bally Ribbon Mills, Collier Corporation, Genesis Engineering, Janicki Industries, Lockheed Martin and Northrop Grumman.

The composite module was fabricated at ATK's Iuka, Miss., facility.

"One of the primary project objectives was to gain hands-on experience for NASA with our contract partners by designing, building and testing a full scale complex structure such as this, then communicate lessons learned to engineers working composites across the agency," said Kirsch.

NESC sponsored the three-year CCM project as part of its mission to solve technical problems related to spaceflight and to make spaceflight safer. The CCM is an all-composite representation of the part-metal, part-composite flight crew module Orion, which is part of NASA's Constellation Program to return man to the moon and/or Mars.


Composite Crew Module (CCM) Pressure Vessel Pathfinder Development



A quote from Robert Zubrin's book Entering Space: Creating a Spacefaring Civilization brought to mind a key advantage of this reconfigured X-33/VentureStar that I hadn't considered before:

"The shuttle is a fiscal disaster not because it is reusable, but
because both its technical and programmatic bases are incorrect. The
shuttle is a partially reusable launch vehicle: Its lower stages are
expendable or semi-salvageable while the upper stage (the orbiter ) is
reusable. As aesthetically pleasing as this configuration may appear
to some, from an engineering point of view this is precisely the
opposite of the correct way to design a partially reusable launch
system. Instead, the lower stages should be reusable and the upper
stage expendable. Why? Becasue the lower stages of a multi-staged
booster are far more massive than the upper stage: so if only one or
the other is to be reusable, you save much more money by reusing the
lower stage. Furthermore, it is much easier to make the lower stage
reusable, since it does not fly as high or as fast, and thus takes
much less of a beating during reentry. Finally the negative payload
impact of adding those systems required for reusability is much less
if they are put on the lower stage than the upper. In a typical two-
stage to orbit system for example every kilogram of extra dry mass
added to the lower stage reduces the payload delivered to orbit by
about 0.1 kilograms, whereas a kilogram of extra dry mass on the upper
stage causes a full kilogram of payload loss. The Shuttle is actually
a 100-tonne to orbit booster, but because the upper stage is a
reusable orbiter vehicle with a dry mass of 80 tonnes, only 20 tonnes
of payload is actually delivered to orbit. From the amount of smoke,
fire, and thrust the Shuttle produces on the launch pad, it should
deliver five times the payload to orbit of a Titan IV, but because it
must launch the orbiter to space as well as the payload, its net
delivery capability only equals that of the Titan. There is no need
for 60-odd tonnes of wings, landing gear and thermal protection
systems in Earth orbit, but the shuttle drags them up there (at a cost
of $10 million per tonne) anyway each time it flies. In short the
Space Shuttle is so inefficient because it is built upside down."
{emphasis in the original.}
Entering Space, p. 29.

Zubrin makes a key point about that dry weight of 80,000 kg of the orbiter, which is essentially an upper stage, that needs to be carried along to bring that approx. 20,000 kg of payload to orbit. That 4 to 1 ratio of the upper stage dry weight to payload weight struck me because the upper stage for rockets is usually a quite lightweight structure. Then the shuttle is quite poor on this measurement. I then thought of the reconfigured kerosene version of the VentureStar I was considering and realized that it was actually quite good on this scale. It could carry ca. 125,000 payload to orbit with a vehicle dry weight ca. 82,000 kg.
Actually the total shuttle system as a whole is even worse on this scale. This site gives the specifications for some launch systems:

Space Launch Report Library.

Here's the page for the shuttle system:

Space Launch Report: Space Shuttle.

You can calculate the total dry weight by subtracting off the propellant weight from the gross weight for each component. I calculate a total dry weight of 310,850 kg to a payload weight of 24,400 kg, a ratio of 12.7 to 1. In contrast the reconfigured VentureStar has this ratio going in the other direction, that is, the payload weight is larger than the vehicle dry weight.
This is important because the total dry weight is a key parameter for the cost of a launch system. I looked at some of the vehicles listed on the Spacelaunchreport.com page and all the ones I looked had the total dry weight higher than the payload weight. For instance for the Delta IV, it's a dry weight of 37,780 kg to a payload weight of 8,450 kg, for a ratio of 4.5 to 1.
For the Atlas V 401 it was 25,660 kg dry weight to a 12,500 kg payload weight, for a ratio of 2 to 1. This was actually one of the better ones. All the ones I looked at, all had a total dry weight significantly higher than the payload weight, usually at least by a 3 to 4 to 1 ratio.
Then the reconfigured VentureStar would be important in that it could reverse this trend (perhaps for the first time?) in making the dry weight actually less than the payload that could be lofted to orbit. Note that not even the original, planned VentureStar could accomplish this, having a dry weight of about 100,000 kg to a payload capacity of 20,000, a ratio of 5 to 1.
The reconfigured kerosene fueled VentureStar would have a greater propellant mass using dense propellants, but the propellant costs are a relatively small proportion of the launch costs. The more important parameter of dry weight would actually be less.
Note also that the reconfigured kerosene VentureStar could accomplish this feat of having a higher payload capacity than its dry weight, while having a payload capacity that would be close to or exceed that of all the former or planned U.S. launchers, and while being of significantly smaller dimensions. See the attached image drawn to scale showing some key U.S. launchers compared to the VentureStar. Note that despite its small size it would be carrying more payload than the shuttle, the Ares I, the Saturn V and nearly that of the Ares V.
Another factor that I somehow missed when I first wrote on this was the great reduction in launch costs. I somehow didn't make the connection between the increase in payload capacity over the original VentureStar configuration and that of the kerosene fueled one.
The development costs for the VentureStar or any launch vehicle are figured into the launch costs. Then the estimated per kg launch costs of ca. $1,000/kg for the original VentureStar are based on the late 90's estimated development costs for the VentureStar. However, a big part of that development cost was due to the composite design which was significantly more expensive then than now. Recall the point I made before about the reduction in costs of composite materials leading to auto makers including them more and more in passenger cars, and this reduction in cost makes them now economically feasible for an all composite SSTO.
Also, hydrogen engines and associated systems are generally more expensive than kerosene ones. So the reconfigured VentureStar would have a lower cost on that component as well. Then the total development cost even including inflation for the reconfigured VentureStar might be at or even below that of the 1990's estimates for the original hydrogen-fueled VentureStar. This means the per launch costs of the new version should be at or below that of the original version.
But the reconfigured VentureStar can carry 6 times the payload of the original VentureStar! This means the per kg launch costs would be 1/6th as much or only $166 per kg! This is such an extreme reduction in launch costs over the current ones that the calculation I made for how much you could reduce the weight of the propellant tanks has to be done in a more serious fashion.
Note that all the other systems for the VentureStar were progressing well. It was only the relatively trivial problem of not using a strong enough glue for the composite propellant tanks, that led to the program being canceled. Then this is so trivial compared to the complexity of the other systems and the importance of having a fully reusable launch system is so clear, that a better course would have been to open up a competition to find ways of getting the composite tanks to work.
I gave a few different possibilities for lightweighting the propellant tanks in section II of the first post in this thread. A few were theoretical, not being tried before. However, the one involving partitioned tanks is just basic engineering so I'll present a detailed calculation for that in the next post.

Bob Clark



exoscientist":3an7q46v said:
Table of Contents.
II.)Lightweight propellant tanks.
...Still another method might be to model the tanks standing vertically as conical but with a flat front and back, and rounded sides. Then the problem with the front and back naturally trying to balloon out to a circular cross section might be solved by having supporting flat panels at regular intervals within the interior. The X-33 composite tanks did have support arches to help prevent the tanks from ballooning but these only went partially the way through into the interior. You might get stronger a result by having these panels go all the way through to the other side.
These would partition the tanks into portions. This could still work if you had separate fuel lines, pressurizing gas lines, etc. for each of these partitions and each got used in turn sequentially. A preliminary calculation based on the deflection of flat plates under pressure shows with the tank made of aluminum alloy and allowing deflection of the flat front and back to be only of millimeters that the support panels might add only 10% to 20% to the weight of the tanks, while getting similar propellant mass to tank mass ratio as cylindrical tank. See this page for an online calculator of the deflection of flat plates:
eFunda: Plate Calculator -- Simply supported rectangular plate with uniformly distributed loading.
http://www.efunda.com/formulae/solid_me ... niform.cfm
Note you might not need to have a partitioned tank, with separate fuel lines, etc., if the panels had openings to allow the fuel to pass through. These would look analogous to the wing spars in aircraft wings that allow fuel to pass through. You might have the panels be in a honeycomb form for high strength at lightweight that still allowed the fuel to flow through the tank. Or you might have separate beams with a spaces between them instead of solid panels that allowed the fuel to pass through between the beams.

We'll view the X-33 hydrogen tanks standing vertically as conical with flattened front and back. This report on page 19 by the PDF file page numbering gives the dimensions of the X-33 hydrogen tanks as 28.5 feet long, 20 feet wide and 14 feet high:

Final Report of the X-33 Liquid Hydrogen Tank Test Investigation Team.

Call it 9 meters long, 6 meters wide, and 4.3 meters deep for this calculation. I'll simplify the calculation by approximating the shape as rectangular, i.e., uniformly 6 meters wide. See the attached image. Note that the rounded portions of the sides, top and bottom will be considered separately. I'll call the vertical length of each section x, and the bulkhead thickness h. Since the length of the tank is 9m, the number of sections is 9/x.
I'm doing the calculation for kerosene/LOX propellant tanks, but approx. same size as the X-33 tanks. Typically these are pressurized in the 20-40 psi range. I'll take it as 30 psi; call it 2 bar, 2x10^5 Pa. Referring to the drawing of the tank, each bulkhead takes part in supporting the internal pressure of the two sections on either side of it. This means for each section the internal pressure is supported by one-half of each bulkhead on either side of it, which is equivalent to saying each bulkhead supports the internal pressure of one section.
The force on each section is the cross-sectional area times the internal pressure, so 6m*x*(2*10^5 Pa), with x as in the diagram the vertical length of each section. The bulkhead cross-sectional area is 6m*h, with h the thickness of the bulkheads. Then the pressure the bulkheads have to withstand is 6m*x*(2*10^5 Pa)/6m*h = (2*10^5 Pa)*x/h.
The volume of each bulkhead is 6m*h*4.3m. The density of aluminum-lithium alloy is somewhat less than aluminum, call it 2,600 kg/m^3. So the mass of each bulkhead is (2,600 kg/m^3)*6m*h*4.3m = 67,080*h. Then the total mass of all the 9/x bulkheads is (9/x)*67080*h = 603,720*(h/x).
Note that additionally to the horizontal bulkheads shown there will be vertical bulkheads on the sides. These will have less than 1/10 the mass of the horizontal bulkheads because the length of each section x will be small compared to the width of 6m, and will have likewise small contribution to the support of the internal pressure.
The tensile strength of some high strength aluminum-lithium alloys can reach 700 MPa, 7*10^8 Pa. Then the pressure the bulkheads are subjected to has to be less than or equal to this: (2*10^5 Pa)*x/h <= 7*10^8 Pa, so x/h <= 3,500, and h/x => 1/3,500. Therefore the total mass of the bulkheads = 603,720*(h/x) => 172.5 kg. Note we have not said yet how thick the bulkheads have to be only that their total mass is at or above 172.5 kg, for one of the twin rear tanks. It's twin would also require 172.5 kg in bulkhead mass. The third, forward, tank had about 2/3rds the volume of these twin rear tanks so I'll estimate the bulkhead mass it will require as 2/3rds of 172.5 kg, 115 kg. Then the total bulkhead mass would be 460 kg, about 15% of the 3,070 kg tank mass I calculated for the reconfigured X-33.
This is for the bulkheads resisting the outwards pressure of the sections. Notice I did not calculate the pressure inside the tank on the bulkheads from the propellant on either side. This is because the pressure will be equalized on either side of the bulkheads. However, we will have to be concerned about the pressure on the rounded right and left sides of the tank, and the rounded top and bottom of the tank, where the pressure is not equalized on the outside of the tank.
Before we get to that, remember the purpose of partitioning the tank was to minimize the bowing out of the front and back sides from the internal pressure. Consider this page then that calculates the deflection of a flat plat under a uniform load:

eFunda: Plate Calculator -- Clamped rectangular plate with uniformly distributed loading.

"This calculator computes the maximum displacement and stress of a clamped (fixed) rectangular plate under a uniformly distributed load."
http://www.efunda.com/formulae/solid_me ... niform.cfm

In the data input boxes, we'll put 200 kPa for the uniform load, 6 meters for the horizontal distance, .3 m, say, for the vertical distance, and 6 mm for the thickness of the plate. For the vertical distance x I'm taking a value proportionally small compared to the tank width, but which won't result in an inordinate number of partitioned sections of the tank. For the thickness I'm taking a value at 1/1000th the width of the tank, which is common for cylindrical tanks. For the material specifications for aluminum-lithium we can take the Young's modulus as 90 GPa. Then the calculator gives the deflection as only 2.35mm, probably adequate.
However, we still have to consider what happens to the rounded sides and the bottom and top. Look at the last figure on this page:

Thin-Walled Pressure Vessels.

http://www.efunda.com/formulae/solid_me ... vessel.cfm

It shows the calculation for the hoop stress of a cylindrical pressure vessel. The calculation given is 2*s*t*dx = p*2*r*dx, using s for the hoop stress. This implies, s = p*r/t, or equivalently t = p*r/s. So for a given material strength s, the thickness will depend only on the radius and internal pressure.
However, what's key here is the same argument will apply in the figure if one of the sides shown is flat, instead of curved. Therefore in our scenario, the rounded sides, top and bottom, which we regard as half-cylinders, will only need the thickness corresponding to a cylinder of their same diameter, i.e., one of a diameter of 4.3m.
So the rounded portions actually require a smaller thickness than what would be needed for a cylinder of diameter of the full 6m width of the tank.
This means the partitioned tank requires material of somewhat less mass than a cylindrical tank of dimension the full width of the tank plus about 15% of that mass as bulkheads.

Bob Clark

tank drawing


EarthlingX":ojyr23oz said:
Can this help you ?

ORBITER is a free flight simulator

I have not done a proper underhood check of this one, plan to, but if you are interested, i would be willing to help you a bit ..

It probably could. From what I've heard it can calculate velocities and altitudes for launch vehicles given some starting conditions.
I need those to see what would be the actual delta-V achieved by my reconfigured vehicles.

Bob Clark


In the first post of this thread I calculated that switching to kerosene would allow the hydrogen-fueled suborbital X-33 to now become an orbital craft. However, I thought it would be able to carry minimal payload if any.
However, I realize I used too low a value for the density of chilled LOX at 1,160 kg/m^3. It should be actually about 10% higher than the usual 1,142 kg/m^3.

This is described here:

Alternate Propellants for SSTO Launchers.
Dr. Bruce Dunn
Adapted from a Presentation at:
Space Access 96
Phoenix Arizona
April 25 - 27, 1996

In table 2 it gives the densities of some chilled fuels including kerosene, i.e., RP-1, and of LOX. The density given for the chilled kerosene is 867 kg/m^3, and for LOX 1,262 kg/m^3. So for the 296 m^3 volume I was taking for the X-33 propellant tanks and a 2.7 mixture ratio for the NK-33 engine, this gives a kero/LOX propellant mass of 332,600 kg.
Now taking the average Isp of the NK-33 as 315 s, this gives a delta-V for the 21,700 kg dry mass, reconfigured X-33 of 8,797 m/s. But when you take into account you get a 462 m/s velocity boost for free from launching at the equator, you only need about 8,500 m/s delta-V to be provided by the rocket to reach orbit.
This allows us to add payload. Adding 2,300 kg payload, the delta-V becomes 8,500 m/s, sufficient for orbit. We can actually get higher payload than this by using more energetic hydrocarbons than kerosene. For instance in table 2 of Dunn's paper on alternate SSTO propellants, he gives the payload for chilled methylacetylene/LOX as 24% higher than for chilled kero/LOX. This would be a payload of 2,850 kg.
These payload amounts would also allow the X-33 to carry a 2 man crew in its 5 by 10 foot payload bay in a tandem arrangement a la the F-14 seating arrangement.
So you could get a fully reusable, SSTO vehicle at much reduced price than the full-sized VentureStar. This article gives the price to build a new X-33 as $360 million in 1998 dollars:

Adventure star.
http://www.flightglobal.com/pdfarchive/ ... 03141.html

Even taking into account inflation, the cost to build the kerosene-fueled version should be comparable or perhaps even less because of the drop in prices for carbon composites and because kerosene engines are generally cheaper than hydrogen ones.
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.
By Dennis R. Jenkins
http://books.google.com/books?id=DUkl5b ... &q=&f=true

Say the builder expected 25% profit over costs of the vehicle over 100 flights. That would be a charge of $4.5 million per flight. At a 2,850 kg payload capacity that would be $1,580 per kilo, or $720 per pound, to orbit. Not as good as the full-sized VentureStar but still significantly better than current launch prices.

Note that the other half-scale suborbital demonstrators for the NASA RLV program by Rockwell and McDonnell-Douglas could be built for comparable prices and would likewise become full orbital craft by switching to kerosene or other dense propellant. Then we could have 3 separate designs for fully reusable SSTO vehicles at costs that could allow fully private financing that would significantly reduce launch costs and would allow manned flights.

Successful operation of these X-33-sized orbital vehicles at a profit would encourage private financing to build the full-scale VentureStar-sized RLV's that could bring launch costs down to the $100 to $200 per kilo range.

Bob Clark



The SpaceLaunchReport.com site operated by Ed Kyle provides the
specifications of some launch vehicles. Here's the page for the Falcon

Space Launch Report: SpaceX Falcon Data Sheet.

Quite interesting is that the total mass and dry mass values for the
Falcon 1 first stage with Merlin 1C engine give a mass ratio of about
20 to 1. This is notable because a 20 to 1 mass ratio is the value
usually given for a kerosene-fueled vehicle to be SSTO. However, this
is for the engine having high vacuum Isp ca. 350 s. The Merlin 1C with
a vacuum Isp of 304 s probably wouldn't work.
However, there are some high performance Russian kerosene engines that
could work. Some possibilities:

Engine Model: RD-120M.


Engine Model: RD-0234-HC.

However, I don't know if this third one was actually built, being a
modification of another engine that burned aerozine.

Some other possibilities can be found on the Astronautix site:


And on this list of Russian rocket engines:

Russian/Ukrainian space-rocket and missile liquid-propellant engines.
http://www.b14643.de/Spacerockets_1/Div ... ngines.htm

The problem is the engine has to have good Isp as well as a good T/W
ratio for this SSTO application. There are some engines listed that
even have a vacuum Isp above 360 s. However, these generally are the
small engines used for example as reaction control thrusters in orbit
and usually have poor T/W ratios.
For the required delta-V I'll use the fact that a dense propellant
vehicle may only require a delta-V of 8,900 m/s, compared to a
hydrogen-fueled vehicle which may require in the range of 9,100 to
9,200 m/s. The reason for this is explained here:

Hydrogen delta-V.
http://yarchive.net/space/rocket/fuels/ ... eltav.html

Then when you add on the fact that launching near the equator gives
you 462 m/s for free from the Earth's rotation, we can take the
required delta-V that has to be supplied by the kerosene-fueled
vehicle as 8,500 m/s.
I'll focus on the RD-0124 because of its high Isp, 359 s vacuum and
331 s sea level. On the "Russian/Ukrainian space-rocket and missile
liquid-propellant engines" page its sea level thrust is given as
253,200 N, 25,840 kgf. However, the Falcon 1 first stage weighs 28,553
kg. So we'll need two of them. Each weighs 480 kg, so two would be 960
kg. This is 300 kg more than the single Merlin 1C. So the dry mass of
the Falcon 1 first stage is raised to 1,751 kg. There is a RD-0124M
listed on the Astronautix page that only weighs 360 kg, but its sea
level Isp and thrust are not given, so we'll use the RD-0124 until
further info on the RD-0124M is available.
Taking the midpoint value of the Isp as 345 s we get a delta-V of
345*9.8ln(1 + 27102/1751) = 9,474 m/s (!) Note also the achieved delta-
V would actually be higher than this because the trajectory averaged
Isp is closer to the vacuum value since the rocket spends most of the
time at altitude.
This calculation did not include the nose cone fairing weight of 136
kg. However, the dry mass for the first stage probably includes the
interstage weight, which is not listed, since this remains behind with
the first stage when the second stage fires. Note then that the
interstage would be removed for the SSTO application. From looking at
the images of the Falcon 1, the size of the cylindrical interstage in
comparison to the conical nose cone fairing suggests the interstage
should weigh more. So I'll keep the dry mass as 1,751 kg.
Now considering that we only need 8,500 m/s delta-V we can add 636 kg
of payload. But this is even higher than the payload capacity of the
two stage Falcon 1!
We saw that the thrust value of the RD-0124 is not much smaller than
the gross weight of the Falcon 1 first stage. So we can get a vehicle
capable of being lifted by a single RD-0124 by reducing the propellant
somewhat, say by 25%. This reduces the dry weight now since one
RD-0124 weighs less than a Merlin 1C and the tank mass would also be
reduced 25%. Using an analogous calculation as before, the payload
capacity of this SSTO would be in the range of 500 kg.
We can perform a similar analysis on the Falcon 1e first stage that
uses the upgraded Merlin 1C+ engine. Assuming the T/W ratio of the
Merlin 1C+ is the same as that of the Merlin 1C, the mass of the two
of the RD-124's would now be only 100 kg more than the Merlin 1C+.
The dry mass and total mass numbers on the SpaceLaunchReport page for
the Falcon 1e are estimated. But accepting these values we would be
able to get a payload in the range of 1,800 kg. This is again higher
than the payload capacity of the original two stage Falcon 1e. In fact
it could place into orbit the 1-man Mercury capsule.
The launch cost of the Falcon 1, Falcon 1e is only about $8 million -
$9 million. So we could have the first stage for that amount or
perhaps less since we don't need the engines which make up the bulk of
the cost. How much could we buy the Russian engines for? This article
says the much higher thrust RD-180 cost $10 million:

From Russia, With 1 Million Pounds of Thrust.
Why the workhorse RD-180 may be the future of US rocketry.
Issue 9.12 | Dec 2001
"This engine cost $10 million and produces almost 1 million pounds of
thrust. You can't do that with an American-made engine."

This report gives the price of the also much higher thrust AJ26-60,
derived from the Russian NK-43, as $4 milliion:

A Study of Air Launch Methods for RLVs.
Marti Sarigul-Klijn, Ph.D. and Nesrin Sarigul-Klijn, Ph.D.
AIAA 2001-4619
"The main engine is currently proposed as the 3,260
lb. RP-LOX Aerojet AJ26-60, which is the former
Russian NK-43 engine. Thrust to weight of 122 to
1 compares to the Space Shuttle Main Engine’s
(SSME) 67 to 1 and specific impulse (Isp = 348.3
seconds vacuum) is 50 to 60 seconds better than
the Atlas II, Delta II, or Delta III RP-LOX engines.
A total of 831 engines have been tested for
194,000 seconds. These engines are available for
$4 million each, which is about 10% the cost of a
http://mae.ucdavis.edu/faculty/sarigul/ ... 1-4619.pdf

Then the much lower thrust RD-0124 could quite likely be purchased
for less than $4 million. So the single RD-0124 powered SSTO could be
purchased for less than $12 million.

Even though the mathematics says it should be possible, and has been
for decades, it is still commonly believed that SSTO performance with
chemical propulsion is not possible even among experts in the space

Space Tourism is a Hoax
By Fredrick Engstrom and Heinz Pfeffer
11/16/09 09:02 AM ET
"In 1903, the Russian scientist Konstantin Tsiolkovsky established the
so-called rocket equation, which calculates the initial mass of a
rocket needed to put a certain payload into orbit, given that the
orbital speed is fixed at 28,000 kilometers per hour, and that the
maximum speed of the gas exhausted from the rocket that propels it
forward is also fixed.
"You quickly find that the structure and the tanks needed to contain
the fuel are so heavy that you will never be able to orbit a
significant payload with a single-stage rocket. Thus, it is necessary
to use several rocket stages that are dumped on the way up to get any
net mass, i.e. payload, into orbit.
"Let us look at the most successful rocket on the market — the
European Ariane 5. Its start weight is 750 tons, of which 650 tons are
fuel, 80 tons are structure and around 20 tons are left for low Earth
orbit payload.
"You can have a different number of stages, and you can look for minor
improvements, but you can never get around the fact that you need big
machines that are staged to reach orbital speed. Not much has happened
in propulsion in a fundamental sense since Wernher von Braun’s Saturn
rocket. And there is nothing on the horizon, if you discount
controlling gravity or some exotic technology like that. In any case,
it is not for tomorrow."
http://www.spacenews.com/commentaries/0 ... -hoax.html

The Cold Equations Of Spaceflight.
by Jeffrey F. Bell
Honolulu HI (SPX) Sep 09, 2005
"Why isn't Mike Griffin pulling out the blueprints for X-30/NASP, DC-X/
Delta Clipper, or X-33/VentureStar? Billions of dollars were spent on
these programs before they were cancelled. Why aren't we using all
that research to design a cheap, reusable, Single-Stage-To-Orbit
vehicle that operates just like an airplane and doesn't fall in the
ocean after one flight?"
"The answer to this question is: All of these vehicles were fantasy
projects. They violated basic laws of physics and engineering. They
were impossible with current technology, or any technology we can
afford to develop on the timescale and budgets available to NASA. They
were doomed attempts to avoid the Cold Equations of Spaceflight."

Then it is important that such a SSTO vehicle be produced even if
first expendable to remove the psychological barrier that it can not
be done. Once it is seen that it can be done, and in fact how easily
and cheaply it can be done, then there it will be seen that in fact
the production of SSTO vehicles are really no more difficult than
those of multistage vehicles.
Then will be opened the floodgates to reusable SSTO vehicles, and low
cost passenger space access as commonplace as trans-oceanic air

Bob Clark


exoscientist":1zph9ycj said:
Then it is important that such a SSTO vehicle be produced even if
first expendable to remove the psychological barrier that it can not
be done. Once it is seen that it can be done, and in fact how easily
and cheaply it can be done, then there it will be seen that in fact
the production of SSTO vehicles are really no more difficult than
those of multistage vehicles.
Then will be opened the floodgates to reusable SSTO vehicles, and low
cost passenger space access as commonplace as trans-oceanic air

Bob Clark

This last may seem a bit extreme. However, Burt Rutan in this recent video interview about 8 minutes in noted that if rockets were operated with the efficiencies of airliners, the price for a passenger to orbit would be in the $12,000 range:

Big Think Interview With Burt Rutan.
A conversation with the aerospace engineer and founder of Scaled Composites.
March 3, 2010

Robert Zubrin in his book,Entering Space: Creating a Spacefaring Civilization comes to a similar conclusion:

"Current -day rockets, such as the kerosene/oxygen-fueled Atlas can deliver about 1 percent of their takeoff mass to orbit - most (about 90 percent) of the remaining mass is propellant. The cost of a kerosene/oxygen propellant mxture (at 3:1 oxygen/kerosene mixture ratio) is about $0.20/kg. Since the propellant consumed during launch has 90 times the mass of the payload delivered, the propellant cost of sending a mass to orbit is about $18/kg. Assuming a total system operating cost of six times the propellant cost (about double the total cost/fuel ratio of airlines), the resulting price of a rocket ride to orbit would be in the neighborhood of $100/kg, or $10,000 for a 100-kg passenger. There is no fundamental reason why space-launch prices in this range cannot be achieved."
Entering Space, by Robert Zubrin, p.22-23.

Then it is notable that trans-Pacific flights are in this cost range. For instance I earlier did a search on the Japan Airlines site for round trip business class tickets from my town of Philadelphia to Tokyo. It ranged from $6,600 to $21,000:

Select Your Flights

Philadelphia to Tokyo Thursday, April 9, 2009
Tokyo to Philadelphia Tuesday, April 14, 2009
Travelers: 1
Travel class: Business and First
Select your fare: Price differences within a fare type may be due to flight connections or availability.
Prices are per adult passenger and include Taxes and Surcharges.
Fare type Fare description Lowest price
Business Saver Special Restricted. Bed-style seating on most long-haul routes -
Executive Class. more details $6,672.48
Business Saver Restricted. Bed-style seating on most long-haul routes -
Executive Class. more details $7,611.48
Business Normal Flexible. Bed-style seating on most long-haul routes -
Executive Class. more details $12,330.48
First Normal Flexible. World-renowned service and comfort - First Class.
more details $21,589.48

And first class tickets even one-way on Qantas from Los Angeles to Australia are in this price range:

Your Search
From: Los Angeles
To: Sydney
Search Options:
Must travel on these dates
Flexible with dates
From: Los Angeles
To: Sydney
Depart Arrive Morning Afternoon Evening
Stops: Non-stop
All flights
Step 2 - Select your flight
Price displayed is for 1 adult and includes surcharges, fees and taxes.
> View lowest fare around this date
Flights Out: Los Angeles to Sydney - Wed 24 Mar 10
From To Flight Business First
23:50 Los Angeles 08:25
(Fri) Sydney QF108 $7926 $15881
Duration: 14h 35m
22:30 Los Angeles 07:25
(Fri) Sydney QF12 $7926 $15881
Duration: 14h 55m

Then recall for the full-sized VentureStar RLV by switching to kerosene we could get in the range of $166/kg launch costs or $16,600 for a 100 kg passenger.

Bob Clark


exoscientist":fys2o90f said:
Even though the mathematics says it should be possible, and has been
for decades, it is still commonly believed that SSTO performance with
chemical propulsion is not possible even among experts in the space
Then it is important that such a SSTO vehicle be produced even if
first expendable to remove the psychological barrier that it can not
be done. Once it is seen that it can be done, and in fact how easily
and cheaply it can be done, then there it will be seen that in fact
the production of SSTO vehicles are really no more difficult than
those of multistage vehicles.
Then will be opened the floodgates to reusable SSTO vehicles, and low
cost passenger space access as commonplace as trans-oceanic air

In another post I argued that the idea frequently written about in early science fiction of millionaires producing their own manned orbital craft may soon be approaching:

Old time stories about the millionaire rocket developer?

However, the calculation of the Falcon 1 first stage with more efficient engines having SSTO capability leads me to a surprising conclusion: it won't even have to be millionaires who could have such craft.
For instance to "own", in the sense of live in, a million dollar home you don't have to have a million dollar income or even a million dollar net worth.
You just have to make the mortgage payments, which per year can be a fraction of the million dollar cost of the home. This is in the salary range of many just upper class individuals. Then such orbital rockets with financing will be in the cost range of many such individuals.
A combination of factors suggest this is possible. First, with mass production the cost of the rocket structure and of the engines will drop significantly. Also, though the Falcon 1 is priced at about $8 million, remember a large proportion of this is to cover development cost. The large majority of this cost was for the development of the engines. But neither of these two SpaceX engines would be used for the SSTO purpose. Instead would be used the much cheaper for their size Russian engines.
We can estimate how much they would cost based on their size and the costs for much larger, i.e., more powerful, Russian engines. The 1,000,000 lbs. thrust RD-180 costs $10 million. The 400,000 lbs. thrust NK-33 costs $4 million. Based on this we can estimate the cost of the 60,000 lbs. thrust RD-0124 I was recommending for the SSTO purpose as $600,000.
Another reason for why such SSTO's will be lower cost than the $8 million Falcon 1 is that the manufacturing cost is actually only a fraction of the launch cost. See for example the estimates in Tables 3 and 4 here:

When Physics, Economics, and Reality Collide. The Challenge of Cheap Orbital Access.
by John M. Jurist, M.D., Sam Dinkin, Ph.D, David Livingston, DBA
http://www.colonyfund.com/Reading/paper ... o.html#elv

Note then the methods for achieving such high mass ratios as with the Falcon 1 first stage don't appear to be especially hard. See for example the description of the Falcon 1 propellant tanks given here:

Falcon 1 Overview.

They appear to use a combination of methods known for decades such as a common bulkhead and an isogrid design. So using these methods, similar high mass ratios could easily be achieved by other aerospace companies. Actually little in research and development costs would be required for the structures.
There is another key cost that figures into launch costs mentioned in the "When Physics, Economics, and Reality Collide: The Challenge of Cheap Orbital Access" article. That is the cost of range access, usually provided by governments. With wide numbers of privately owned rockets launching daily this cost could be reduced significantly.
A big component of the research and development costs however would be the actual flight tests. This would be significantly reduced with reusable systems.
For this to have a high demand you would need for it to be manned-flight capable. The Falcon 1e sized SSTO would require two RD-0124 engines for a 1,800 kg payload capacity, but would be able to loft a Project Mercury-sized capsule:


You would also need a lightweight reentry system. Inflatable heat shields may fit the bill:

NASA Launches New Technology: An Inflatable Heat Shield.
http://www.nasa.gov/topics/aeronautics/ ... /irve.html

Pod People.
They're the ones thinking outside the space capsule.
* By James Oberg
* Air & Space Magazine, November 01, 2003
"IN 1964, MOST VIEWERS OF TELEVISED SPACE "SHOTS," AS THEY WERE CALLED THEN, knew what it took to protect a spacecraft from the fire of reentry. It took big, heavy shields bolted to pressurized metal vessels. One of the most nerve-racking moments of the early space program had been the final minutes of John Glenn’s 1962 Mercury flight, when Mission Control waited to learn whether his shield had remained attached to the Friendship 7 capsule during the violent return.
"Two years later, on June 10, 1964, another, much lighter vehicle entered the atmosphere with no one on board. In engineering terms it was nearly as daring as the Mercury flights had been. Launched on a sounding rocket to an altitude of 96 miles over New Mexico, the craft dove back toward Earth at a speed of more than 5,000 mph. Being so light, it didn’t generate as much heat from atmospheric friction as Glenn’s capsule had, so it had only a thin coating of thermal protection—no shield. Odder still, it was inflated like a balloon in a Thanksgiving day parade."
http://www.airspacemag.com/space-explor ... oberg.html

This second, which involves a lifting body inflatable heat shield, would result in significant reduction in reentry heating by making use of shapes optimized for high lift/drag ratio at hypersonic speed.

Bob Clark


I'm looking for a numerical trajectory integration program if anyone has access to one.
These are the conditions under which I want to estimate the required delta-V to orbit:

1.)use a dense propellant such as kerosene/LOX; dense propellants are known to reduce gravity losses.

2.)use a moderate to high liftoff thrust/weight ratio, say, 1.4 and above; high liftoff T/W also reduces gravity losses.

3.)launch near equator to get the ca. 460 m/s tangential boost.

4.)only get to 100 km, the altitude considered space, to just launch satellites or make orbital transfers, not for long term orbits.

Bob Clark


On this forum and in email correspondence the objection has been made to
developing the SSTO that it would not be cost effective because two-stage-to-
orbit systems could lift more payload.
But in addition to dry mass being a key indicator of cost another key parameter of
cost is complexity of the system. Single stages are inherently simpler than
multi-stages so a SSTO would win on that scale.
Moreover if the vehicle were to be a fully reusable two-stage system, then
you would have the cost and complexity of making two separate winged vehicles.
However, the SSTO can actually help if you wanted to get a TSTO to loft
larger payloads. For instance I argued that the X-33, when kerosene fueled,
could become orbital, though with small payloads, and the full-sized
VentureStar could loft large payloads when kerosene fueled. However, the
full-sized VentureStar would be quite expensive, in the billions of dollars
range, while the X-33 would be about $360 million to build a new one. So the
smaller X-33 could launch smaller payloads at an smaller initial investment.
But quite key also is that using two of them as the first and second stages,
they could now serve as a heavy launch system, and this would be at a much
smaller investment than building the full-size VentureStar.
Note also the comparison to a two-stage expendable system: the two stages in
such a case would be expendable because they don't have sufficient mass ratio
to singly get to orbit. That is they aren't weight optimized. But suppose you
were able to make each stage be so optimized that each separately could reach
orbit at the same size vehicles. Then now note this means these weight
optimized versions could therefore loft *more* payload because the weight
savings could go to extra payload AND would be less costly per launch in being
This then would be a key advantage of making a small reusable orbital
launcher as I maintain the X-33 can be when kerosene fueled. For your early
forms of two-stage to orbit expendable systems, you might want to save on
development costs, and just get it done easily by not optimizing it for
weight. But IF at some later time you do have the technological development to
so weight optimize the stages so they can each be reusable and separately
orbital, then you make better BOTH your payload and your costs to orbit.

Bob Clark


Another question asked in email and on different forums about this is, if the Russians already had these high performance kerosene engines that make SSTO possible why aren't they doing it?
I wondered about that too. I thought their new Angara rocket should be SSTO capable since it will be using these high performance kerosene engines. But then I checked the specifications of the Angara rocket on the SpaceLaunchReport.com site:

Space Launch Report: New Launchers - Angara.

I found that the Angara mass ratios are significantly worse than for the SpaceX Falcon launchers. In fact, in general the Russian launchers are not as well mass optimized as the American launchers:

Is It Safe?
* By Michael Milstein
* Air & Space Magazine, May 01, 2009
"Russian spacecraft, says NASA spokesman John Yembrick, rely heavily on beefier mechanical structures for safety rather than complex backup systems. In the mid-1990s, NASA compared the design and standards for the Russian Soyuz spacecraft to its own and concluded that both NASA and Roscosmos, Russia's space agency, have equivalent safety requirements, though the Russians follow a different path to meet those parameters. NASA's decision to put American astronauts on Soyuz for a ride to the space station was based on the rocket's history of safety and reliability."
http://www.airspacemag.com/space-explor ... c=y&page=2

This probably is a big part of the reason the Russians have had this great drive to increase the performance of their kerosene engines - out of necessity.
Then to get SSTO you use the best features of both the American and Russian designs combined into one.

Bob Clark


According to SpaceX they have a vacuum version of the Merlin engine with a vacuum Isp of 342 s:

New Merlin Vacuum engine demonstrates highest efficiency for an American hydrocarbon rocket engine.
"McGregor, TX. – March 10, 2009 – Space Exploration Technologies (SpaceX) successfully conducted a full mission duration firing of its new Merlin Vacuum engine on March 7, at SpaceX's Test Facility in McGregor, Texas. The engine fired for a full six minutes, consuming 100,000 pounds of liquid oxygen and rocket grade kerosene propellant.
"The new engine, which powers the upper stage of SpaceX's Falcon 9 launch vehicle, demonstrated a vacuum specific impulse of 342 seconds – the highest efficiency ever for an American hydrocarbon rocket engine. Thrust was measured at approximately 92,500 lb of force in vacuum conditions and the engine remained thermally stable over the entire run."

To have this efficiency while being optimized for vacuum operation a rocket engine has to have a longer and heavier nozzle. This greatly reduces its efficiency at sea level. However, a version of this vacuum Merlin that uses instead a aerospike nozzle might be able to maintain its high efficiency at sea level as well.
This might also make the engine lighter but I've seen conflicting reports on whether using an aerospike nozzle really makes an engine lighter.
Though not yet used on actual orbital launch vehicles, the aerospike has been flight tested:

http://en.wikipedia.org/wiki/Aerospike_ ... erformance

According to these refs the aerospike gets high efficiency from sea level to high altitude conditions:

Aerospikes (Henry Spencer; Ben Muniz; Jeff Greason)

Nozzle Design.
by R.A. O'Leary and J. E. Beck
Threshold Magazine, Spring 1992

They do mention there is some diminution in efficiency in the velocity range of Mach 1 to Mach 3, but overall its performance is well above that of the usual bell nozzle.
I had suggested that a Falcon 1 first stage that used a high performance Russian engine in stead of the Merlin 1C could make a SSTO. But of course SpaceX itself would be disinclined to use a different engine from the one they have spent millions of dollars developing. But if an aerospike Merlin could maintain correspondingly high performance at sea level, then SpaceX could have an SSTO from an aerospike Merlin Falcon 1 first stage as well.

Bob Clark
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