The Coming SSTO's.

[Orionblamblam]

Single-stage-to-orbit.
Examples.
http://en.wikipedia.org/wiki/Single_stage_to_orbit#Examples

WEAK.

The first stage of the Titan II had the mass ratio required for single-stage-to-orbit capability with a small payload. A rocket stage is not a complete launch vehicle, but this demonstrates that an expendable SSTO was probably achievable with 1962 technology.

Have you run trade studies showing that an expendable Titan II first stage derived SSTO with its miniscule payload provide a real economic advantage over other launch vehicles? How much would a T-II S1 SSTO cost? How much would just the bare T-II S1 cost? How much would it cost to design the needed changes? What would they cost to produce? What would they cost to operate? How much would the new nosecone weigh? How about the control systems? The ACS systems? The new engines?

 

[Michel Van]

let face it, SSTO have some major Disadvantages

the biggies is: very low payload with high launchmass, if the SSTO is to reused
and if you raise the Payload, the Launchmass gets ridiculously BIG

of course you find a solution in High tech
like tanks from Litium-aluminum,composite material
and take Aerospike engine or even exotic airbreathing rocket engine
but all this is VERY Expensively in R&D and build cost, i mean billions and billions of US Dollar, Pound, Euros

next to that is the preparation process for the Launch, Mission, landing and Overhaul for the next mission.the Space Shuttle neede 500 employees, who take mounths to this job
but employees need to be payed, Wat raise the programme cost
and how more complex the SSTO (see High Tech), more employees are needed. Wat raise the Cost higher
upss Wat not that problem the STS programme run into ?

So was is the Ideal solution ?
Two Stage To Orbit with only First stageto reused, build from existing hardware !
no extrem R&D cost, higher payload as SSTOwhy on reuse the first stage only ?
most biggies part in a Rocket is the first stage also put most fuel into energy to launch the second stage.
best example of low cost TSTO is Aerojet Seadragon
http://www.secretprojects.co.uk/forum/index.php/topic,12874.0.html

 

[RGClark]

let face it, SSTO have some major Disadvantages
the biggies is: very low payload with high launchmass, if the SSTO is to reused
and if you raise the Payload, the Launchmass gets ridiculously BIG
...

The key point is that the payload does not have to be very small. It could be quite large in relation to a key figure of merit, as indicated by the example in the first post in this thread.
As I said in that first post comparing the payload to the gross mass of the vehicle is actually a poor way to measure the efficiency of a launch system. The reason is you would be comparing the payload to something mostly made up of what contributes minimally to the cost of the launch, which is the cost of the propellant.
A much better comparison is the payload mass to the dry mass of the vehicle, since this is what actually has to be constructed and tested. Experts in the industry are becoming more and more aware of this fact. See for instance this report:

A Comparative Analysis of Single-Stage-To-Orbit Rocket and Air-Breathing Vehicles.
http://govwin.com/knowledge/comparative-analysis-singlestagetoorbit-rocket-and/15354

From the report:

From the many vehicle parameters used in the design of an RLV, a few parameters were assumed to be the most significant and used as figures of merit in this study. In both aircraft and spacecraft design, a vehicle’s empty mass is used as a guide to predict the vehicle’s design, materials, manufacturing, quality control and operational costs [2, 21]. Smaller vehicle empty mass is considered favorable.

p. 5.

Compared to the cost of the RLV, the cost of fuel is relatively insignificant [5]. Vehicle gross mass, consisting mostly of mass due to fuel, was therefore not considered to be a major figure of merit in this study.

p. 5.

4.2 Empty Mass Trends
The gross takeoff mass and empty mass are plotted for all vehicles in this study in Figure 20. Gross mass does not indicate where mass is allocated (structure, payload or mass) and consists of mostly inexpensive propellant. It is presented here for reference. Empty mass is a good indication of procurement and operational costs because it consists of the expensive structure of the vehicle.

p. 52.

and

Empty weight is considered a good figure of merit for the total cost of procuring a vehicle and one of the main figures of merit for maintaining and operating a vehicle.

p. 67.

Now note that the example I gave had a payload mass to dry mass ratio greater than 1. For every other orbital rocket I looked at, and probably for any orbital rocket that has ever existed, this ratio is going in the other direction: the total dry mass for all stages is greater than the payload mass. And often it is much greater. For the space shuttle system for example the total dry mass is 12 times that of the payload. This probably is a key reason why the space shuttle is such an expensive system for the payload delivered, despite being partially reusable.


Bob Clark

 

[Matej]

I must say I like Scott's arguments much than others. Non-reusable SSTO must be less effective than the TSTO by principle - everything is packed together, meaning higher weight during the flight profile = at least higher fuel consumption with the same engines and thus higher costs. You also need some sort of adaptive propulsion to compensate the density of the air (something like experimental Aerospike).

Regarding the reusable SSTO, the main driving factor for their development was high flight rate and "airplane like" operating scheme to compensate its disadvantages. But "airplane like" doesn't mean "exactly like". Even if you are able somehow to build a reusable SSTO and make a lot of sorties per time period, it will require extensive post-flight check to make it operational for the next flight. Note that post-flight check of the reusable spaceplane (accelerates in minutes to the 10 km/s, operates in microgravity/electromagnetic smog from the sun/cosmic dust, must sustain temperatures from nearly absolute zero to hundreds of degrees Celsia during the reentry) is something significantly different compared to check done on the airplane flying from London to Paris. If you want to do another compensation and use the robust construction, it usually means higher weight, minimizing usefull cargo and bigger purchasing costs, so you are once again at the beginning of the problem.

I must say I like those cool looking SSTO vehicle concepts, but I don't find them effective. The best way, how to do effective payload delivery into space is (in my opinion) the conventional rocket with the reusable airbreathing first stages and expendable second stage. And for the human spaceflights, lets hope for the significant progress in the material research to allow construction of the wire, strong enough to support the space elevator.

 

[ouroboros]

Fuel costs are a red herring though since they are still such a low percentage of the overall costs for the foreseeable future. Driving costs right now are parts count (on a generic basis). That's why Sea Dragon is so attractive, as the parts count is on the same order as conventional rockets, but only the material costs have increased due to larger total mass in comparison to achieved payload. Though Sea Dragon is a TSTO though...


The beamed power rocket guys make a decent case for SSTO using a laser or microwave absorber heat exchanger surface heating a hydrogen fuel (no oxidizer) rocket, but all too often leave out the important part of getting the vehicle high enough for the beaming site to get a fix on the vehicle in more practical site configurations. Which rather quickly gets back to TSTO. About the only solution I've seen to the "First 5Km" problem that wasn't some kind of conventional first stage rocket derivative or a carrier aircraft that looked practical is a Quicklaunch style immersed tube in the ocean lined with a linear rail track. Sling the vehicle up the tube (evacuated or not, you don't want to break the sound barrier at the muzzle anyways) to toss it within line of sight of the beaming station. Using tube immersion tricks you could potentially go up to 1Km or more above the ocean at the muzzle. If you wanted to go the evacuated tube route, you would be looking at 8MWe/m of tube diameter at the muzzle for a plasma valve/window to seal the tube.

Beamed rockets like those proposed by Kare are in the 800-1000 ISP range, and that assumes no tricks like an air-turbo rocket or something like a LANTRN afterburner or even an aerospike. Though Kare's baseline expendable SSTO design looks like an airfoil cross section shape with an expendable pop top shell and internal twin drop tanks. The absorber plate could in theory be used for reentry protection in a reusable SSTO if you circulate cold hydrogen ullage from the main fuel tank.

The guys at Escape Dynamics prefer microwave systems, but if the military research on high power diode lasers, or even the guys at Lasermotive get the unit laser size down to the envisioned container sized level, laser will probably happen before microwave as a viable beam array source.

And like Mr. Lowther says, any other research that improves the viability of SSTO/TSTO helps the beamed power rocket guys too...

 

[Orionblamblam]

Beamed-power systems have the *potential* to be true game changers... *if* they can be made to work. And it's by no means certain that they can anytime soon. But if they can, in principle the spacecraft would carry a small quantity of a cheap consumable such as liquid hydrogen, liquid nitrogen or even just water, and would expend a whole lot of electricity. The costs I've seen for the electricity are on the order of half a dollar per pound of payload to orbit. But just as fuel costs are a small fraction of the cost of a chemical rocket... running that laser array just might be more expensive than imagined.

But *if* that can all be worked out, beamed power systems are the only systems I've come across that at least have the potential of lowering space launch costs to the point where average people could actually dream of going to orbit.

 

[sublight]

Hey Scott what about an SSTO using a combo of rockets and a Gerald Bull type cannon? Any idea how many G's the average payload could accept? (Assuming one could do a "soft start" cannon blast)

 

[RyanC]

Darryl Davis, who leads Boeing's Phantom Works, tells AvWeek that they are proposing a 3-4 year technology readiness assessment that would lead up to a demonstration of a X-37B type of system but would be smaller. Wind tunnel tests have been completed. Davis says the system would be a single stage capable of reaching low Earth orbit and, with a booster, higher orbits. The system would return to Earth as a glider. Davis says "that advances in lightweight composites warrant another look" at single-stage-to-orbit launchers.

What I think the USAF wants is something that can be used to rapidly launch tactical satellites (50-100 lb) into very low orbits which would decay in a few months to provide rapid battlefield surveillance and communications on demand, with a growth option to insert heavier payloads into more normal orbits via adding a flyback booster stage.

 

[Orionblamblam]

Hey Scott what about an SSTO using a combo of rockets and a Gerald Bull type cannon? Any idea how many G's the average payload could accept? (Assuming one could do a "soft start" cannon blast)

Cannon launch might be fine for extremely rugged microsats and densely-packed commodites (ingots of steel or aluminum, say, maybe even water), but not for much of anything else. Cannon have launched solid-state electronic payloads at up to 20,000 gees or so (such as the Copperhead anti-tank shell and atomic cannon rounds), but I'm not sure if a solir rocket circularization motor could survive that . The thign to do would probably be to put a closure on the aft end of the nozzle and fill the void volume in the rocket motor with a liquid/metal slurry as dense or denser than the solid propellant; this would cushion the propellant from the acceleration loads and could be easily jettisoned.

Still, the payload potential of such a system would be pretty minimal.

 

[RGClark]

Titan 2 first stage:
Inert mass 9000 lb
Gross mass 269,000 lbs
Vacuum Isp 296 sec --> Delta V= 9860 m/sec
Sea level Isp 258 --> Delta V=8590 m/sec

Dr. John Schilling provides a launch vehicle performance estimator on his company's web site based on a numerical formula:

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

Using the numbers for the Titan II and a value for the vacuum thrust of 2,108.5 kN I've seen on other sites for the first stage and launching from Cape Canaveral at an inclination of 28.5 degrees to an altitude of 185 km gives this result:

Mission Performance:
Launch Vehicle: User-Defined Launch Vehicle
Launch Site: Cape Canaveral / KSC
Destination Orbit: 185 x 185 km, 28 deg
Estimated Payload: 1507 kg
95% Confidence Interval: 419 - 2844 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.

I used Schilling's launch estimator also on the Atlas-derived SSTO I discuss in the first post and got a payload over 4,000 kg using the estimator, in the range of what I calculated in that first post.


Bob Clark

 

[RanulfC]

So was is the Ideal solution ?
Two Stage To Orbit with only First stageto reused, build from existing hardware !
no extrem R&D cost, higher payload as SSTOwhy on reuse the first stage only ?
most biggies part in a Rocket is the first stage also put most fuel into energy to launch the second stage.
best example of low cost TSTO is Aerojet Seadragon
http://www.secretprojects.co.uk/forum/index.php/topic,12874.0.html

Actually as I understand it studies tend to show that as backwards, upper-stages tend to cost more than lower stages. Though those same studies tend to point to "simple, cheap, and robust" (read also as less efficent, but larger and tougher) "booster" stages that can be recovered, but can also be expendable.
(Which as I understand it is a variable economics argument that depends on things like flight rate, manufacturing costs, and refubishment costs. Low flight rates tend towards expendable operations, and "high" flight rates tend to be MUCH higher than anything the current market would need)

As far as I can see this tends to run the suggested TSTO designs into what seem to be "Assisted-SSTO" Orbiters coupled to BDB-Boosters. Which actually is an "argument" used by SSTO advocates I notice, if the SSTO "almost" gets to orbit then all you need to do is "add-a-booster" (or boosters) like adding GEMs to an expendable for the needed performance boost.

Still, that IS a "TSTO" I guess so OBB's argument makes sense, except I get the feeling from his posts that his interpritation might be a bit different? Would the oft repeated phrase of "fully-reusable-TSTO" cover such vehicles?

Randy

 

[Orionblamblam]

Using the same online payload estimator, and data from Isakowitz for the T-II stage 1:

dry mass: 4000
Propellant: 118000
Thrust: 2090
Isp: 278 (an average of sea level& vacuum)
Payload fairing: 500 kg jettisoned at 200 seconds
Cape Canaveral launch to 185 X 185 km, 28 degree inclination
Direct ascent:

Result:
Estimated payload: zero kg
95% confidence interval: 0 - 419 kg
-----------------------
Running the same thing again, but with an Isp of 288:
Result:
Estimated payload: 0 kg
95% coinfidence interval: 0 - 1036 kg

The 500 kg mass for the payload fairing is less than the actual 652 kg Titan II payload fairing.

Not included are the neccesary additional masses ofan attitude contro,k system and associated avionics.

So... no. Titan II first stage doesn't seem to be able to make it to orbit on its own. It would require a weight-saving program as well as more efficient engines.

 

[Orionblamblam]

Actually as I understand it studies tend to show that as backwards, upper-stages tend to cost more than lower stages.

Yup. The way the math works is that for a 1 kg dry-mass increase for an SSTO, you get a 1 kg reduction in payload. But for a TSTO, a 1 kg dry mass increase in stage 1 might result in a 0.1 kg decrease in payload; a 1kg dry mass increase in stage 2 might mean a 0.9 kg decrease in payload. Consequently, first stages tend to be more ruggedly buily - and thus more easily and cheaply built - than upper stages. Additionally, every launch vehicle needs an expensive control system... a control system that stays with the vehicle all the way to orbit. This means it's located in the upper stage, not the lower one. While the lower stage will certainly have gimbalable nozzles, throttleable engines, attitude control thrusters, jet vanes, whatever... the actual *commands* will come from the second stage. So the upper stage will not only require tighter tolerances, it'll also require vastly more computational smarts.

 

[RGClark]

Using the same online payload estimator, and data from Isakowitz for the T-II stage 1:
dry mass: 4000
Propellant: 118000
Thrust: 2090
Isp: 278 (an average of sea level& vacuum)
Payload fairing: 500 kg jettisoned at 200 seconds
Cape Canaveral launch to 185 X 185 km, 28 degree inclination
Direct ascent:
Result:
Estimated payload: zero kg
95% confidence interval: 0 - 419 kg
-----------------------
Running the same thing again, but with an Isp of 288:
Result:
Estimated payload: 0 kg
95% coinfidence interval: 0 - 1036 kg
The 500 kg mass for the payload fairing is less than the actual 652 kg Titan II payload fairing.
Not included are the neccesary additional masses ofan attitude contro,k system and associated avionics.
So... no. Titan II first stage doesn't seem to be able to make it to orbit on its own. It would require a weight-saving program as well as more efficient engines.

I was puzzled about which Isp value to use for the calculator so I emailed Dr. Schilling about it and he said you should always use the vacuum value. He said the program takes into account the reduction in Isp and thrust at sea level and low altitude.
Using a vacuum value of 296 s and a payload fairing of 500 kg jettisoned at 200 s gives this result:

Launch Vehicle: User-Defined Launch Vehicle
Launch Site: Cape Canaveral / KSC
Destination Orbit: 185 x 185 km, 28 deg
Estimated Payload: 1007 kg
95% Confidence Interval: 0 - 2340 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

 

[Orionblamblam]

Estimated Payload: 1007 kg
95% Confidence Interval: 0 - 2340 kg

So, still zero, and possibly negative. Note that the "estimated payload" falls on the low end of the "95% confidence interval," meaning that a negative payload is still possible even with best-case assumptions.

 

[RGClark]

Estimated Payload: 1007 kg
95% Confidence Interval: 0 - 2340 kg

So, still zero, and possibly negative. Note that the "estimated payload" falls on the low end of the "95% confidence interval," meaning that a negative payload is still possible even with best-case assumptions.

True the Schilling calculator is only an estimate. Its best *estimate* is that it would be 1,000 kg. It could possibly be 0 or it could be as high as 2,300 kg. The way to know for sure would be to run one of the accurate programs NASA uses such as POST that calculates the delta-V over an actual trajectory track.


Bob Clark

 

[Orionblamblam]

True the Schilling calculator is only an estimate. Its best *estimate* is that it would be 1,000 kg. It could possibly be 0 or it could be as his high as 2,300 kg.

No. It could be as high as 2300 kg. It could be as low as -300 kg.

The way to know for sure would be to run one of the accurate programs NASA uses such as POST that calculates the delta-V over an actual trajectory track.

It would only make sense to do that *after* you've actually defined the vehicle. So far, you have only defined *part* of the vehicle. How much do the avionics/power/control systems weight? How about the attitude control systems? How about the modifications needed to the engines for stop/start? How about the Isp hit that the engine will take when you institute a deep-throttling requirement?

So far you haven't even gotten to "back of the envelope" level of accuracy yet.

 

[Michel Van]

That Titan II example remind me of the SSTO Saturn V S-ID proposal (Aka Saturn V-B)

back in 1967 Boeing Proposed to modified the Saturn V first stage S-IC
into a 1/2 Stage like the Atlas ICBM
means drop 4xF-1 with part of Stage Thrust structure at 154 sec after liftoff
and remaining F-1 bring the rest with payload into orbit


so with Wat we get ?
launch mass: 2300000 kg or 5099000 lb
Payload: 22600 kg or 49800 lb
So a SSTO who launch mass is 101 higher as the Payload...

the Saturn IB its only 28 heavier as its Payload of 22600 kg.

 

[RGClark]

True the Schilling calculator is only an estimate. Its best *estimate* is that it would be 1,000 kg. It could possibly be 0 or it could be as his high as 2,300 kg.

No. It could be as high as 2300 kg. It could be as low as -300 kg.

The way to know for sure would be to run one of the accurate programs NASA uses such as POST that calculates the delta-V over an actual trajectory track.

It would only make sense to do that *after* you've actually defined the vehicle. So far, you have only defined *part* of the vehicle. How much do the avionics/power/control systems weight? How about the attitude control systems? How about the modifications needed to the engines for stop/start? How about the Isp hit that the engine will take when you institute a deep-throttling requirement?
So far you haven't even gotten to "back of the envelope" level of accuracy yet.

The statements that the Titan II first stage had SSTO capability though with minimal payload goes back several years, well before for example when Schilling came up with his online calculator. I gather there were done launch simulations with POST for example.


Bob Clark

 

[Orionblamblam]

So, I'll just wait patiently for you to produce those weights and performace changes, yes?

 

[RanulfC]

Here's the Saturn (and Shuttle-ET) article:
http://www.spacefuture.com/archive/a_single_stage_to_orbit_thought_experiment.shtml

Someone on NSF did the math:
http://forum.nasaspaceflight.com/index.php?topic=23703.msg678883#msg678883

According to that post the Titan-II doesn't have enough delta-v total, but it turns out he was mistaken on his figures, the updated figures are here:
http://forum.nasaspaceflight.com/index.php?topic=23703.msg679069#msg679069

However, Mitchell Burnside Clapp points out that due to propellant density among other things he thinks the Titan-II first stage WOULD work as an expendable SSTO:
http://groups.google.com/group/sci.space.policy/browse_thread/thread/3d981607d59684dc/945baea33c95a22?q=group:sci.space.*+author:Burnside-Clapp+PFSMF&fwc=1

He also "re-designed" a NASA studied LOX-LH2 SSTO into a LOX-Kero SSTO:
http://www.erps.org/papers/LO2-KeroseneSSTO.html

More discussion on "Dense Propellant" applications
http://yarchive.net/space/rocket/fuels/hydrogen_deltav.html

Not that "high-density" propellant SSTO hasn't been proposed before:
http://www.quantumg.net/mockingbird.pdf
http://www.dunnspace.com/alternate_ssto_propellants.htm
http://www.sworld.com.au/steven/pub/nsto.pdf
http://www.osti.gov/bridge/servlets/purl/379977-2LwFyZ/webviewable/379977.pdf
http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19770003210_1977003210.pdf
http://www.erps.org/papers/SSTORwNCP.pdf

From what I'm reading THE major issue with the Titan-II first stage SSTO is that the engines actually produce far to much thrust to fly an efficent trajectory.
However most of these while they might "prove" SSTO don't have very much payload, (There are a couple of exceptions) and very aggresive mass/fractions to close the design(s).

You can increase payload by staging, but how much does that scale up/down anyway?

Take for example the High-Density-Propellant SSTO suggested by Mitchell Clapp in the last cited paper:
"Alternative Single Stage to Orbit Design
An altemative single stage to orbit design vehicle was designed using hydrogen peroxide and JP-5 as the propellants. A sketch of the vehicle appears in figure 3. The emphasis through out the design effort was efficient operations, easy maintenance, and simplicity. The mission of the vehicle was to place a 10,000 pound payload of 15 ft diameter and 20 ft length into a polar orbit.

A vehicle meeting this specification would be capable of orbiting a 23,500 pound payload by means of a due east launch from White Sands Missile Range, New Mexico.

The payload sits in its bay at the approximate midpoint of the vehicle. The fuel tank is in the nose and the oxidizer tank is aft of the payload. The payload is located where it is to minimize travel of the center of gravity when the payload is removed.
The vehicle is aerodynamically stable in nose first or tail first reentry and can achieve an lift to drag ratio of 0.8, sufficient to give it aerodynamic cross range of 600 miles. The thermal protection system weighs 1.2 pounds per square foot, on average, and is composed of carborn/silicon carbide, metal multiwall, and thermal blanket materials. The vehicle lands tail first on engine thrust only.

The engines are pressure fed and operate at a chamber pressure of 14 MPa (2,000p psi) There are thirty engines, arranged in a circle around the base of the vehicle. Each engine has an exparrsion ratio of 18.5, but the cluster as a whole has an expansion ratio of 366.

This yields a specific impulse at launch of 289 s, and in vacuum a value of 335 s. The zero-length-plug nozzle configuration causes some lose of efficiency ( 4.5 percent ) but the reduction in weight is favorable.

Thrust vectoring is accomplished by individual gimbiling of the engines. The tanks are composed of the highest strength-to-weight material available, Kevlar-49. For all structural components, a weight margin of l5 percent was retained along with standard safery factors for pressure vessels and buckling. The vehicle is designed with a takeoff thrust to weight of 1.3, and can lose up to six engines before aborting. Secure intact abort even with multiple engine failures is a key opertational requirement for an effctive and safe launcher.

The veliicle is 56 ft long and 23 feet in diameter at the base. The empty mass including payload, is 30,000 pounds. The gross mass is 620,000 pounds. It is designed for a crew of two and an endurance on orbit, of four days."

I suspect the weight margins are a bit thin, but suppose the number DO work and it looks like it can make orbit on it's own. Which is cheaper to retain the payload? A full up second reusable stage or strap on boosters?

Even if it's an "almost" SSTO the simplicity and operational characteristics would seem to make for a lower over all launch cost. Or am I wrong somewhere?

Randy

 

[RGClark]

The point of the matter is that the many small spacecraft and suborbital craft of lightweight composite design become high Mach suborbital, a la the X-33, when switched to using high efficiency engines. And moreover if they are scaled up by a factor of 2, then these larger versions become fully orbital vehicles.
...
This is also true of the X-34 and SpaceShipOne: they become high Mach suborbital, as a single stage, when switched to high efficiency engines. And when scaled up twice as large with the high efficiency engines, they become now fully orbital single stage vehicles.
The case of SpaceShipOne is especially interesting because the twice scaled up vehicle is already built in SpaceShipTwo. Then swapping out the hybrid engines of SpaceShipTwo for high efficiency ones produces a SSTO.

SpaceShipeOne is given a dry mass of 1,200 kg:

SpaceShipeOne.
http://en.wikipedia.org/wiki/Spaceshipone

We'll fill the entire fuselage aft of the pilot's cabin up until the nozzle with kerosene/LOX propellant. I'll estimate dimensions from the image attached below. The cylindrical portion of the fuselage is about 10 feet long. After this there is a tapered portion of the fuselage that extends up to the nozzle, about 7.5 feet long. The cylindrical portion is about 5 feet wide. The narrow end of the tapered portion is about 1.5 feet wide.
The tapered portion is in the shape of a frustum:

Volume of a Frustum of a Cone.
http://jwilson.coe.uga.edu/emt725/Frustum/Frustum.cone.html

By the volume formula on that page, it's volume will be (1/3)*Pi*(7.5)(2.5^2 + 2.5*.75 + .75^2) = 68.2 cu. ft.
The volume of the cylindrical portion of the fuselage will be Pi*10*(2.5)^2 = 196.34 cu. ft., for a total of 264.57 cu. ft., or 7.5 cubic meters. The overall density of kerolox is about 1,000 kg/m^3. So this will have about 7,500 kg of propellant.
We need to replace the hybrid engine and tanks with kerolox engines and tanks. Astronautix gives the hybrid engine of SpaceShipOne a mass of 300 kg:

SpaceDev Hybrid.
http://www.astronautix.com/engines/spaybrid.htm

Removing this gives the engine-less SpaceShipOne a mass of 900 kg. For a replacement kerolox engine we'll use the RD-0242-HC at 120 kg:

RD-0242-HC.
http://www.friends-partners.org/partners/mwade/engines/rd0242hc.htm

As I mentioned before the high chamber pressure suggests this is a high performance engine. With altitude compensation it should get a vacuum Isp in the range of 360 s. As a point of comparison the rather low efficiency Merlin 1C just by using a longer, vacuum optimized nozzle increases its vacuum Isp from 305 s to 342 s. The high efficiency Russian engines also can get a sea level Isp in the range of 331 s. So we'll take this as the sea level Isp using altitude compensation. Using the estimate of Ed Kyle of the trajectory averaged Isp being 2/3rds of the way from the sea level value to the vacuum value, we'll take the average Isp as 350 s.
We also have to add the mass of the kerolox tanks. Their mass will be about 1/100th that of the mass of propellant so at 75 kg. The total dry mass will now be 1,095 kg.
To this we add thermal protection. The advanced ceramics used on the Air Force's X-37B mass about 12 kg/m^2. The cross-sectional area to be covered on the bottom of the vehicle from the tip of the nose cone to the end of the tapered section of the fuselage will be about 8.5 square meters. This gives a thermal protection mass of 102 kg for the fuselage. For the wings, the wing area is 15 m^2 resulting in a thermal protection mass of 180 kg for the wings. So the total mass is now 1,377 kg, call it 1,380 kg.
Then the delta-V will be 350*9.8ln(1 + 7500/1380) = 6,385 m/s. Even adding the total mass of two pilots at 200 kg, the delta-V would still be 5,998 m/s.
So it will be a high Mach suborbital craft. As I'll show in a following post the twice scaled up SpaceShipTwo will be a fully orbital craft when switched out to use high efficiency liquid fueled engines.
At this high a delta-V though this reconfigured SpaceShipOne could also be used for the Air Force's Reusable Booster System program. This is intended to cut launch costs by using a reusable booster and an expendable upper stage. This will give a small low cost proof of principle version of the system.


Bob Clark

 

[RGClark]

...
More discussion on "Dense Propellant" applications
http://yarchive.net/space/rocket/fuels/hydrogen_deltav.html

Not that "high-density" propellant SSTO hasn't been proposed before:
http://www.quantumg.net/mockingbird.pdf
http://www.dunnspace.com/alternate_ssto_propellants.htm
http://www.sworld.com.au/steven/pub/nsto.pdf
http://www.osti.gov/bridge/servlets/purl/379977-2LwFyZ/webviewable/379977.pdf
http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19770003210_1977003210.pdf
http://www.erps.org/papers/SSTORwNCP.pdf
...

Thanks for those links. I had seen those other reports appearing in the 1990's suggesting dense propellants would make it easier to make a SSTO but I had not seen before this one you cited:

HIGH DENSITY PROPELLANTS FOR SINGLE STAGE TO ORBIT VEHICLES.
http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19770003210_1977003210.pdf

That this appeared in the 1970's is interesting because it is my thesis that SSTO's carrying significant payload became possible in the 1970's with the advent of the high performance kerosene and hydrogen fueled engines created then:

SSTO's would have made possible Arthur C. Clarke's vision of 2001.
http://www.secretprojects.co.uk/forum/index.php/topic,13212.0.html

Indeed I argue these SSTO's could have made been low cost enough to be privately owned, of comparable cost to medium sized business jets. However, these small, low cost ones, which would have made passenger flight to space routine, would have been kerosene fueled since dense propellants make possible smaller sized SSTO vehicles. Then it is important to know that this concept of dense propellants making possible smaller, cheaper SSTO's was at least known of then.
The problem was it was not generally accepted then, just as was not the possibility of SSTO's generally accepted then. Unfortunate, because if it had, we already would have had by now manned exploration of the solar system, financed by the enormous wealth in minerals available from asteroids. We wouldn't be fretting now about what we are going to do for manned space flight in the U.S. now that the shuttle is retired.
However, the point of the matter is we already have the solution to low cost manned space access now, SSTO's.


Bob Clark

 

[RanulfC]

I'd not seen that one either and probably STILL wouldn't have if I hadn't found the link off hand in an archive message. Funny but it doesn't seem to come up on a normal word search on "dense propellants" but I suspect that's because despite the title the paper is ONLY considering "dense" propellants to be used "mixed-mode" applications. (LOX with an engine that burns both LH2 and Dense-propellants or a mixture of engines)
It DOES come up when using the key words "mixed-mode" or "Tri-Propellant" in the search so that is probably what was used as keywords.

In that context it was most likely produced for use on studies of the Salkeld concepts using mixed mode engines and propellants for SSTO or near-SSTO vehicles:
http://www.astronautix.com/craft/saluttle.htm
http://www.pmview.com/spaceodysseytwo/spacelvs/sld039.htm

Of course the "issue" with the SSTO concepts was even using tri-propellant and best optimization to get what was considered a "useful" payload (25,000lbs or more) to orbit required a huge vehicle. In most studies TSTO or at least 1.5STO came out with better economics and operations than a pure SSTO.

Randy

 

[RGClark]

I'd not seen that one either and probably STILL wouldn't have if I hadn't found the link off hand in an archive message. Funny but it doesn't seem to come up on a normal word search on "dense propellants" but I suspect that's because despite the title the paper is ONLY considering "dense" propellants to be used "mixed-mode" applications. (LOX with an engine that burns both LH2 and Dense-propellants or a mixture of engines)
It DOES come up when using the key words "mixed-mode" or "Tri-Propellant" in the search so that is probably what was used as keywords.
In that context it was most likely produced for use on studies of the Salkeld concepts using mixed mode engines and propellants for SSTO or near-SSTO vehicles:
http://www.astronautix.com/craft/saluttle.htm
http://www.pmview.com/spaceodysseytwo/spacelvs/sld039.htm
Of course the "issue" with the SSTO concepts was even using tri-propellant and best optimization to get what was considered a "useful" payload (25,000lbs or more) to orbit required a huge vehicle. In most studies TSTO or at least 1.5STO came out with better economics and operations than a pure SSTO.
Randy

Yes, that article did seem to focus only on mixed mode or tri-propellant propulsion. This it was argued would increase the payload over just using hydrogen alone. There have been studies on converting high performance Russian kerosene engines to mixed mode as well as the SSME's to mixed mode going back to the 70's. So these might have been implemented then allowing routine space travel by the 2001 time frame.

In regards to getting the most economical delivery of payload to orbit. Quite key here is that if you use the principle of using both the most lightweight stages and the most efficient engines at the same time then you can loft even more payload to orbit with your mult-stage launchers. Plus, the individual stages can now be used as SSTO's to loft smaller payloads at a lower cost than using the full multi-stage launchers.
I mentioned before that SpaceX is using weight optimized design for their Falcon 9 launcher. They are getting a 20 to 1 mass ratio for the Falcon 9 first stage. And they expect to achieve a 30 to 1 mass ratio for the side boosters on their Falcon Heavy. If they had used high efficiency engines such as the NK-33 or the RD-180 instead of the Merlins on their Falcons they could loft even more payload to orbit as well as using the first stages or boosters alone as SSTO's to launch smaller payloads.
It is notable that Elon Musk this week announced that SpaceX will be working on a "super efficient" engine which he says will allow reusable launchers that can bring the price to orbit down to $50 to $100 per pound, in the range of what I was saying. The key point is this is doable now with the high efficiency engines already existing and the lightweight stages already existing.

August 03, 2011
Looking at Spacex plans for Making Falcon Rockets Reusable to get to $50 per pound launch costs.
http://nextbigfuture.com/2011/08/looking-at-spacex-plans-for-making.html

August 02, 2011
Elon Musk of Spacex talks about a Reusable Falcon Heavy to get to $50 a pound to space.

Two technology areas Musk didn’t like were lifting bodies/wings and nuclear rockets.
On the former, he said he was a “vertical takeoff, vertical landing” type guy and eschewed wings since they had to be tailored for each planet’s atmosphere and were useless on airless bodies such as the Moon.
Drawbacks to nuclear power included the need for shielding (heavy), water (heavy), and public objections against launching nuclear fuel on a rocket. “It’s a tricky thing getting a reactor up there with a ton of uranium,” Musk said and went on to say while nuclear power would be useful for Mars or lunar operations, he implied that some assembly (i.e., mining and processing fuel off planet) would be required.

http://nextbigfuture.com/2011/08/elon-musk-of-spacex-talks-about.html

c.f.,

SSTO's would have made possible Arthur C. Clarke's vision of 2001.
http://www.secretprojects.co.uk/forum/index.php/topic,13212.0.html

Bob Clark

 

[RanulfC]

Well they kind of HAVE to get around to optimizing the Merlin engines, for Kero/LOX engines their really pretty bad

Stage recovery seems to be Space-Xs biggest hurdle right now, but they are stuck as they can't add margin without loosing payload UNLESS they make some major changes. (I'll read the link first)

As to Elon's "opinions" I leave them to him. After all it's HIS money he's spending But designing something that "could" be used on other planets to get from the surface-to-LEO-and-back of Earth isn't doing yourself any favors. He may not "like" wings and wheels but they ARE highly optimizable for operations on Earth and THAT is the issue he and everyone else is currently dealing with, NOT landing on the Moon or Mars

Randy

 

[RGClark]

Quite key for why reusable SSTO's will make manned space travel routine is the small size and low cost they can be produced. A manned SSTO can be produced using currently existing engines and stages the size of the smallest of the very light, or personal, jets [1], except it would use rocket engines instead of jet engines, and the entire volume aft of the cockpit would be filled with propellant, i.e., no passenger cabin. So it would have the appearance of a fighter jet.
We'll base it on the SpaceX Falcon 1 first stage. According to the Falcon 1 Users Guide on p.8 [2], the first stage has a dry mass of 3,000 lbs, 1,360 kg, and a usable propellant mass of 47,380 lbs, 21,540 kg. We need to swap out the low efficiency Merlin engine for a high efficiency engine. However, SpaceX has not released the mass for the Merlin engine. We'll estimate it from the information here, [3]. From the given T/W ratio and thrust, I'll take the mass as 650 kg.
We'll replace it with the RD-0242-HC, [4]. This is a proposed modification to kerosene fuel of an existing hypergolic engine. This type of modification where an engine has been modified to run on a different fuel has been done before so it should be doable [5], [6]. The engine mass is listed as 120 kg. We'll need two of them to loft the vehicle. So the engine mass is reduced from that of the Merlin engine mass by 410 kg, and the dry mass of the stage is reduced down to 950 kg. Note that the mass ratio now becomes 23.7 to 1.
We need to get the Isp for this case. For a SSTO you want to use altitude compensation. The vacuum Isp of the RD-0242-HC is listed as 312 s. However, this is for first stage use so it's not optimized for vacuum use. Since the RD-0242-HC is a high performance, i.e., high chamber pressure engine, with altitude compensation it should get similar vacuum Isp as other high performance Russian engines such as the RD-0124 [7] in the range of 360 s. As a point of comparison the Merlin Vacuum is a version of the Merlin 1C optimized for vacuum use with a longer nozzle. This increases its vacuum Isp from 304 s to 342 s [8]. I've also been informed by email that engine performance programs such as Propep [9] give the RD-0242-HC an ideal vacuum Isp of 370 s. So a practical vacuum Isp of 360 s should be reachable using altitude compensation.
For the sea level Isp of the RD-0242-HC, again the version of the high performance, high chamber pressure, RD-0124 with a shortened nozzle optimized for sea level operation gets a 331 s Isp. So I'll take the sea level Isp as this value using altitude compensation that allows optimized performance at all altitudes.
To calculate the delta-V achievable I'll follow the suggestion of Mitchell Burnside Clapp who spent many years designing and working on SSTO projects including stints with the DC-X and X-33 programs. He argues that you
should use the vacuum Isp and just use 30,000 feet per second, about 9,150 m/s, as the required delta-V to orbit for dense propellants [10]. The reason for this is that you can just regard the reduction in Isp at sea level and low altitude as a loss and add onto the required delta-V for orbit this particular loss just like you add on the loss for air drag and gravity loss. Then with a 360 s vacuum Isp we get a delta-V of 360*9.8ln(1 + 21,540/950) = 11,160 m/s. So we can add on payload mass: 360*9.8ln(1+21,540/(950 + 790)) = 9,150 m/s, allowing a payload of 790 kg.
To increase the payload we can use different propellant combinations and use lightweight composites. Dr. Bruce Dunn wrote a report showing the payload that could be delivered using high energy density hydrocarbon fuels other than kerosene [11]. For methylacetylene he gives an ideal vacuum Isp of 391.1 s. High performance engines can get get ca. 97% and above of the ideal Isp so I'll take the vacuum Isp value as 384 s. Dunn notes that Methyacetylene/LOX when densified by subcooling gets a density slightly above that of kerolox, so I'll keep the same propellant mass. Then the payload will be 1,120 kg: 384*9.8ln(1 + 21,540/(950 + 1,120)) = 9,160 m/s.
We can get better payload by reducing the stage weight by using lightweight composites. The stage weight aside from the engines is 710 kg. Using composites can reduce the weight of a stage by about 40%. Then adding back on the engine mass this brings the dry mass to 670 kg. So our payload can be 1,400 kg: 384*9.8ln(1 + 21,540/(670 + 1,400)) = 9,160 m/s.
Note this has a very high value for what is now regarded as a key figure of merit for the efficiency of a launch vehicle: the ratio of the payload to the dry mass. The ratio of the payload to the gross mass is now recognized as not being a good figure of merit for launch vehicles. The reason is that payload mass is being compared then to mostly what makes up only a minor proportion of the cost of a launch vehicle, the cost of propellant. By comparing instead to the dry mass you are comparing to the expensive components of the vehicle, the parts that have to be constructed and tested [12].
This vehicle in fact has the payload to dry mass ratio over 2. Every other launch vehicle I looked at, and possibly every other one that has ever existed, has the ratio going in the other direction, i.e., the dry mass is greater than the payload mass. Often it is much greater. For example for the space shuttle system the dry mass is over 12 times that of the payload mass, undoubtedly contributing to the high cost for the payload delivered.
Because of this high value for this key figure of merit, this vehicle would be useful even as a expendable launcher. However, a SSTO is most useful as a reusable vehicle. This will be envisioned as a vertical take-off vehicle. However, it could use either a winged horizontal landing or a powered vertical landing. This page gives the mass either for wings or propellant for landing as about 10% of the dry, landed mass [13]. It also gives the reentry thermal protection mass as 15% of the landed mass. The landing gear mass is given as 3% of the landed mass here [14]. This gives a total of 28% of the landed mass for reentry/landing systems. With lightweight modern materials quite likely this could be reduced to half that.
If you use the vehicle just for a cargo launcher with cargo left in orbit, then the reentry/landing system mass only has to cover the dry vehicle mass so with lightweight materials perhaps less than 100 kg out of the payload mass has to be taken up by the reentry/landing systems. For a manned launcher with the crew cabin being returned, the reentry/landing systems might amount to 300 kg, leaving 1,100 kg for crew cabin and crew. As a mass estimate for the crew cabin, the single man Mercury capsule only weighed 1,100 kg [15 ]. With modern materials this probably can be reduced to half that.
For the cost, the full two stage Falcon 1 launcher is about $10 million. The engines make up the lion share of the cost for launchers. So probably much less than $5 million just for the 1st stage sans engine. Composites will make this more expensive but probably not much more than twice as expensive. For the engine cost, Russian engines are less expensive than American ones. The RD-180 at 1,000,000 lbs vacuum thrust costs about $10 million [16], and the NK-43 at a 400,000 lbs vacuum thrust costs about $4 million [17]. This is in the range of $10 per pound of vacuum thrust. On that basis we might estimate the cost of the RD-0242-HC of about 30,000 lbs vacuum thrust as $300,000. We need two of them for $600,000.
So we can estimate the cost of the reusable version as significantly less than $10,600,000 without the reentry/landing system costs. These systems added on for reusability at a fraction of the dry mass of the vehicle will likely also add on a fraction on to this cost. Keep in mind also that the majority of the development cost for the two stage Falcon 1 went to development of the engines so in actuality the cost of just the first stage without the engine will be significantly less than half the full $10 million cost of the Falcon 1 launcher. The cost of a single man crew cabin is harder to estimate. It is possible it could cost more than the entire launcher. But it's likely to be less than a few 10's of millions of dollars.

REFERENCES.
1.)List of very light jets.
http://en.wikipedia.org/wiki/List_of_very_light_jets

2.)Falcon 1 Users Guide.
http://www.spacex.com/Falcon1UsersGuide.pdf

3.)Merlin (rocket engine)
4 Merlin 1C Engine specifications
http://en.wikipedia.org/wiki/Merlin_(rocket_engine)#Merlin_1C_Engine_specifications

4.)RD-0242-HC.
http://www.astronautix.com/engines/rd0242hc.htm

5.)LR-87.
http://en.wikipedia.org/wiki/LR-87

6.)Pratt and Whitney Rocketdyne's RS-18 Engine Tested With Liquid Methane.
by Staff Writers
Canoga Park CA (SPX) Sep 03, 2008
http://www.space-travel.com/reports/Pratt_and_Whitney_Rocketdyne_RS_18_Engine_Tested_With_Liquid_Methane_999.html

7.)RD-0124.
http://www.astronautix.com/engines/rd0124.htm

8.)Merlin (rocket engine).
2.5 Merlin Vacuum
http://en.wikipedia.org/wiki/Merlin_(rocket_engine)#Merlin_Vacuum

9.)Propep
http://www.spl.ch/software/index.html

10.)Newsgroups: sci.space.policy
From: Mitchell Burnside Clapp <cla...@plk.af.mil>
Date: 1995/07/19
Subject: Propellant desity, scale, and lightweight structure.
http://groups.google.com/group/sci.space.policy/browse_frm/thread/3d981607d59684dc/945baea33c95a22?hl=en

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

12.)A Comparative Analysis of Single-Stage-To-Orbit Rocket and Air-Breathing Vehicles.
p. 5, 52, and 67.
http://govwin.com/knowledge/comparative-analysis-singlestagetoorbit-rocket-and/15354

13.)Reusable Launch System.
http://en.wikipedia.org/wiki/Reusable_launch_system#Horizontal_landing

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

15.)Mercury Capsule.
http://www.astronautix.com/craft/merpsule.htm

16.)Wired 9.12: From Russia, With 1 Million Pounds of Thrust.
http://www.wired.com/wired/archive/9.12/rd-180.html

17.)A Study of Air Launch Methods for RLVs.
Marti Sarigul-Klijn, Ph.D. and Nesrin Sarigul-Klijn, Ph.D.
AIAA 2001-4619
p.13
http://mae.ucdavis.edu/faculty/sarigul/aiaa2001-4619.pdf

 

[Byeman]

Nothing but a bunch of unsubstantiated assumptions (like all composite structure for the stage). Too many optimizations.

This is what happens when a non engineer tries to design.

 

[RGClark]

Nothing but a bunch of unsubstantiated assumptions (like all composite structure for the stage). Too many optimizations.
This is what happens when a non engineer tries to design.

You don't need to use composites. Just using higher energy fuels than kerosene can also result in high payload. I mentioned methylacetylene, but several others result in high payload as well, which can be confirmed by the several rocket engine performance programs freely available on the net.
But in regard to composites several references state the weight savings can be in this range:

The Policy Origins of the X-33
The DC-XA
December 22, 1999
Part II

Lee also explained to Goldin that primary structures, such as aeroshell sections, made of graphite composite materials would reduce overall structural weight by 40 percent compared to the Shuttle structure, without corrosion or fatigue problems. Actual transport aircraft with composite primary structures included the Boeing 777, the Airbus A330 and A340, and the ATR 72. In addition, Lee wrote, NASA would study both tri-propellant rocket engines and an upgraded Space Shuttle Main Engine (SSME) for use on a single-stage-to-orbit vehicle.

http://www.hq.nasa.gov/pao/History/x-33/facts_62.htm

National Security Space Road Maps (NSSRM).
Structural Systems (U)
Overview (U):

(U) The goal of this subthrust is to enhance the capability and performance of the payload by reducing the mass of necessary structural support and thermal bus susbsystems. The approach to meeting this goal includes the application of advanced composite materials to both satellite and launch vehicle structural components, the research and development of higher heat flux thermal bus components, and the ability to test new structural materials in a realistic simulation of the space environment. The ultimate aim is to satisfy the Air Force's needs for lower cost and lighter weight spacecraft and launch vehicle systems.
...
(U)Launch Vehicle Structures: The objective of this area is to reduce launch vehicle costs and weights by at least 50% and cut lead times by a third through the application of advanced composite materials and structural designs. Light-weight launch vehicle structures manufactured using advanced composite materials enable existing and next-generation space launch systems to carry more payload into orbit. In addition, these structures may be manufactured using techniques that can substantially reduce costs and/or lead times compared to conventional methods. Programs are also being pursued that address vehicle health monitoring, i.e., systems capable of processing data from traditional non-destructive evaluation sources as well as on-board sensors for determining launch readiness.

http://www.fas.org/spp/military/program/nssrm/initiatives/strucsys.htm

Single Stage to Orbit:
A Reliable Transport System or an Unattainable Dream?

Since the dry mass of an SSTO vehicle must be kept low, it requires the use of high-tech materials and technologies. While these were deemed unattainable in early concept designs, many "futuristic" materials are now in common use. Using an aluminum-lithium alloy for fuel tanks instead of pure aluminum can reduce the total vehicle weight by 4%, and allow surrounding materials to be lighter as well, and it actually reduces the total vehicle weight by 23%. Also, the use of composite materials for the vehicle's structure reduces the total vehicle weight by 45% (Bekey 34).

http://vorlon.case.edu/~jam64/work/ssto.htm

Bob Clark

 

[Byeman]

Proof that you don't know what you are talking.

Current launch vehicles already use composites for all non tank or thrust sections structures.

None of your quotes are proof. Show hardware that proves the point.

Your whole MO is to spout disjointed links from doing nothing but internet searches and flood forums with your nonsense. Have you put a design to paper and had the results reviewed by competent parties? Have you assembled anything close to aerospace hardware or even been near some?

 

[RGClark]

Proof that you don't know what you are talking.
Current launch vehicles already use composites for all non tank or thrust sections structures.
None of your quotes are proof. Show hardware that proves the point.
Your whole MO is to spout disjointed links from doing nothing but internet searches and flood forums with your nonsense. Have you put a design to paper and had the results reviewed by competent parties? Have you assembled anything close to aerospace hardware or even been near some?

Some correspondents with NASA have suggested I submit them to NASA open RFI's.

For using composites for getting lightweight reusable vehicles, there is of course the Air Force's X-37B.
And the Air Force is requesting composite design including propellant tanks for its Reusable Booster System (RBS) program:

USAF Seeks Reusable Booster Ideas.
May 14, 2009
By Graham Warwick

AFRL's reference concept includes an integral all-composite airframe and tank structure that carries both internal pressure and external flight loads. The concept vehicle is powered by pump-fed liquid-oxygen/hydrocarbon rocket engines.

http://www.aviationweek.com/aw/generic/story_channel.jsp?channel=space&id=news/Reuse051409.xml

Then Boeing realized that the Air Force requirement of all composite design would actually permit the production of an SSTO:

Boeing proposes SSTO system for AF RBS program.

The new issue of Aviation Week has a brief blurb about a Boeing
proposal for the Air Force's Reusable Booster System (RBS) program:
Boeing Offers AFRL Reusable Booster Proposal - AvWeek - June.13.11
(subscription required).
Darryl Davis, who leads Boeing's Phantom Works, tells AvWeek that they
are proposing a 3-4 year technology readiness assessment that would
lead up to a demonstration of a X-37B type of system but would be
smaller. Wind tunnel tests have been completed. Davis says the system
would be a single stage capable of reaching low Earth orbit and, with
a booster, higher orbits. The system would return to Earth as a
glider.
Davis says "that advances in lightweight composites warrant another
look" at single-stage-to-orbit launchers.

http://www.hobbyspace.com/nucleus/index.php?itemid=30110

Compare this to what I wrote soon after the Air Force announced the RBS program in 2009:

Newsgroups: sci.space.policy, sci.astro, sci.physics
From: Robert Clark <rgregorycl...@yahoo.com>
Date: Sat, 19 Dec 2009 16:35:12 -0800 (PST)
Subject: Re: A kerosene-fueled X-33 as a single stage to orbit vehicle.
...

The Air Force is researching reusable hydrocarbon-fueled first stage
boosters to be used with expendable upper stages to cut the costs to
space by 50%:

USAF Seeks Reusable Booster Ideas.
May 14, 2009
By Graham Warwick
"AFRL's reference concept includes an integral all-composite airframe
and tank structure that carries both internal pressure and external
flight loads. The concept vehicle is powered by pump-fed liquid-oxygen/
hydrocarbon rocket engines."
http://www.aviationweek.com/aw/generic/story_channel.jsp?channel=space&id=news/Reuse051409.xml

This article discusses wind tunnel tests of a scale-model of such a
booster:

AEDC team conducts first test on a reusable space plane.
Posted 12/16/2009 Updated 12/16/2009
http://www.arnold.af.mil/news/story.asp?id=123182588

It is interesting they are proposing an all-composite construction
including propellant tanks for this reusable hydrocarbon-fueled first
stage booster.
As I have argued, an all-composite, hydrocarbon-fueled design would
allow even a reusable single-stage-to-orbit vehicle.

Bob Clark

http://groups.google.com/group/sci.space.policy/msg/f98ee8868109b914?hl=en

Perhaps Boeing decided to do this because of what I wrote on various forums in 2009 and in correspondence with NASA, Air Force, and industry officials then. Or perhaps just in running the numbers for a RBS first stage booster they realized that it could actually be a SSTO. It really is quite obvious once you run the numbers.
The disagreements I've had with people on this topic always stems from not just looking at what the numbers say. It comes from just making the assumption SSTO's can't work.

Perhaps it will take now another two years for people in the industry to recognize that manned SSTO's the size of small business jets are possible. But I doubt it will even take that long. Once you realize that SSTO's are possible the rest follows easily.



Bob Clark

 

[Byeman]

For using composites for getting lightweight reusable vehicles, there is of course the Air Force's X-37B.
And the Air Force is requesting composite design including propellant tanks for its Reusable Booster System (RBS) program:

Then Boeing realized that the Air Force requirement of all composite design would actually permit the production of an SSTO:

You are reading the wrong things into them

a. Use of composites on spacecraft is not something new, it goes back decades. The fact that X-37 is a reusable spacecraft has nothing to do with launch vehicles.

b. Boeing only says to reassess it, it doesn't say it is feasible.

EDITED:

 

[Orionblamblam]

There is a long, long history of people making postings on Usenet and the like where they do some simple calculations and declare that they have Solved All The Problems That Have Plagued The Experts. The sci.space newsgroups have long been bombarded with such from the likes of William Mook and others who crank out tons of verbage and micrograms of actual testing.

 

[RGClark]

There is a long, long history of people making postings on Usenet and the like where they do some simple calculations and declare that they have Solved All The Problems That Have Plagued The Experts. The sci.space newsgroups have long been bombarded with such from the likes of William Mook and others who crank out tons of verbage and micrograms of actual testing.

Again, just run the numbers. Boeing did and reached the same conclusion I did two years ago.

Bob Clark

 

[Orionblamblam]

Again, just run the numbers.

No. BUILD IT.

The world is full to the brim with Brilliant Ideas That Work Great On Paper. It's not so full to the brim with people willing to actually do the work. If you are so certain of your math, I'm sure you can build your own. If your math is ironclad and complete, you can find investors. An SSTO that can carry a meaningful payload and cost on the order of $10,000,000 should have the world beating a path to your doorstep. And as such... yammering on about it on newsgroups and online fora is the entirely wrong use of your time. instead, you should be making presentations to investors.

At the very least, i look forward to your AIAA papers describing your design and your progress. Be sure to post links to the YouTube videos of the launches.

 

[Grey Havoc]

You are reading the wrong things into them

a. Use of composites on spacecraft is not something new, it goes back decades. The fact that X-37 is a reusable spacecraft has nothing to do with launch vehicles.

To be fair, composites technology hasn't matured quite as fast as people would have expected in the late 1980's and '90's, especially when it comes to space applications. For example, while the use of composites in some programs such as the original DCX (in particular the aeroshell) was a success, in other programs such as the X-33/VentureStar (that fuel tank!), they were -AHEM!- problematic.

 

[RGClark]

Again, just run the numbers.

No. BUILD IT.

We are agreed on that.

Bob Clark

 

[Orionblamblam]

We are agreed on that.

So when can we expect to see updates on *your* physical progress? After all, if you are correct, then you shopuld have no problem getting investment capital PDQ and getting design & construction work underway.

 

[RGClark]

It would be a truly watershed moment just creating a SSTO even if it doesn't carry much payload. It wouldn't have to be anything extensive like perhaps what Boeing is planning with their X-37B derived SSTO.
A small one could be demonstrated by amateur science or technical organizations, for instance by the British Interplanetary Society, or the Planetary Society.
The Planetary Society is spending about $5.8 million total on their two attempts at solar sail demonstators:

Cosmos 1.
http://en.wikipedia.org/wiki/Cosmos_1

LightSail-1.
http://en.wikipedia.org/wiki/LightSail-1#Creation

A small SSTO demonstrator that could carry a few hundred pound payload could be developed for less than this amount and would be far more important for it would show that low cost SSTO's are possible.
In fact the organization developing it could even make money on it because they could use it to launch small scientific payloads.

For the purpose of just making the demonstration it might work to make the vehicle half the size of the one I described here:

http://www.secretprojects.co.uk/forum/index.php/topic,13211.msg132292.html#msg132292

So it would use one RD-0242 engine, have a propellant load about 10,000 kg, and, perhaps, have a dry weight of 475 kg. However, vehicle dry weights don't scale linearly. Scaling a vehicle up actually improves your mass ratio. So by making the vehicle half-scale we probably would not get as good a mass ratio, i.e., the dry mass would likely be more than just half that of the full sized vehicle.
In addition to the amateur science organization funded test SSTO's, it might be funded as an X-prize competition. This might have the same effect as the Ansari X-Prize had in spurring commercial suborbital ventures. It would spur manned commercial orbital ventures.
However, these would need high performance turbopump fed engines. This is an entire level of difficulty above that of the suborbital rockets which just use pressure-fed engines. In fact the complexity of turbopump fed engines have led rocket engineers to opine "orbital launchers are turbopump developments with rockets attached".
I recommend teams attempting the venture engage in partnerships with Aerojet or Pratt & Whitney who have experience with high chamber pressure, turbopump-fed engines, especially of the Russian type. They both also have experience in converting an engine from one fuel to another, Aerojet with the conversion of the Titan II engines from kerosene to hypergolics, and Pratt & Whitney more recently with the conversion of the Apollo lunar lander engines from hypergolics to methane.
Their costs would be partially defrayed by the amount of the X-prize. This prize amount should at least be that of the $30 million total prize money offered for the Google Lunar X-Prize competition, since its importance greatly exceeds it. Note too for such prize competitions the amount spent by the teams often exceeds that offered by the prize. They could also be offered a portion of the profits that would come from development of the vehicles as small payload orbital launchers.
For this prototype test vehicle you probably would not need to use the SpaceX weight optimized Falcon 1 first stage since you just want to get positive payload to orbit. Interestingly I found that Armadillo Aerospace has successfully used common bulkhead design which saves significantly on tank weight for their suborbital test rockets. They would be a good choice for a low cost stage.
However, Armadillo has not been successful in their last two suborbital test flights, apparently due to failures in guidance and control. Though Armadillo apparently has solved this for hovering vehicles, it is a significantly more difficult problem for a vehicle traveling at high speed. I recommend a partnership with the MIT Draper labs. They did the G & C for the Apollo missions. More recently they are engaged in partnerships to win the Google Lunar X-Prize.


Bob Clark

 

[RGClark]

A common estimate is that orbital flight is an order of magnitude more difficult than suborbital flight, as measured for example by the energy requirements. On that basis a prize for a commercial manned flight to orbit might be 10 times of the suborbital X-Prize, so to $100 million. This is actually probably doable considering that the required engines and stages already exist to do it as an SSTO.
Another source for funding might be Bigelow Aerospace. Bigelow had offered a prize in 2004 of $50 million for a commercial reusable manned launcher to orbit. The prize though expired in January 2010 with no takers:

America's Space Prize.
http://en.wikipedia.org/wiki/America's_Space_Prize

However, the original Orteig Prize for a non-stop cross Atlantic flight also expired with no takers. It was the second offer of the prize for an additional 5 year period which was won by Lindbergh.
Then Bigelow could offer the manned space flight prize for an additional 5 year period. But the original conditions for the prize were probably too ambitious. Bigelow appeared to want manned transport craft to his Bigelow space hotels to be fully developed from the winner of the prize in his requiring a 5 man vehicle. However, following the example of the suborbital X-Prize, just accomplishing a small 1 man test flight would be sufficient to serve as an impetus for commercial ventures to invest in developing such launchers aside from the prize.
Then I suggest Bigelow lower the requirement to only needing a single crew member. This would allow multiple test flights before a manned flight is attempted.


Bob Clark