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An SSTO as "God and Robert Heinlein intended".

Posted by: RGClark - Tue Jan 04, 2011 8:37 am
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An SSTO as "God and Robert Heinlein intended". 
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Rocket Constructor
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Post Re: An SSTO as "God and Robert Heinlein intended".   Posted on: Sun Jul 10, 2011 12:33 am
http://www.xcor.com/press-releases/2007 ... sting.html

Rocket has been tested. Structures done also


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Post Re: An SSTO as "God and Robert Heinlein intended".   Posted on: Sun Jul 10, 2011 2:55 am
The XCOR engine you refer to is not has no chance of being part of an SSTO attempt.

It is a project, jointly with P&W (I think), and is a demo of a methane engine.

It's a fairly low pressure one, using pressurized system--no turbopump.

Again, for reasons [1] and [2] in previous post apply.


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Post Re: An SSTO as "God and Robert Heinlein intended".   Posted on: Sun Jul 10, 2011 3:58 pm
marsmission2020 wrote:
SSTO refuelable to mars is unique though

Kerosine it would be cold in deeper space

And liquid hydrogen will boil off


So methane ssto is desirable in that context

100000 Lbm fuel. 5000 Lbm rocket and 5000 payload nearly


With the higher Isp of methane over kerosene, that should be doable.

Bob Clark

_________________
Single-stage-to-orbit was already shown possible 50 years ago with the Titan II first stage.
Contrary to popular belief, SSTO's in fact are actually easy. Just use the most efficient engines
and stages at the same time, and the result will automatically be SSTO.
Blog: http://exoscientist.blogspot.com


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Post Re: An SSTO as "God and Robert Heinlein intended".   Posted on: Sun Jul 10, 2011 4:14 pm
ckpooley wrote:
SSTO is fundamentally unworkable. Even thouigh the back-of-envelope amth says yes, just possible it is no good because:
[1] lift-off engine must work at sea level and vacuum. big compromise, and
none of the altitude adapting designs are good enough
[2] aerodynamic loads will require the stage to be heavier and shaped for
withstanding the loads.
So 2 stage makes sense. Have the 1st stage carry the upper stage high enough
for the engine to work as if in vacuum, and enclose the upper stage so it does not contend with aerodynamic loads.
The delta-v for 1st stage need not be much over 1 km/sec for staging to be at 60 km. The Microlauncher entry level launcher, "ML-1", is to have a delta-v of just over 2 km/sec to overcome gravity and aerodynamic loss. The staging altitude is to be 60-70 km, with an upward velocity of 500-800 m/sec.
The upper stage (2 stages in ML case) can then be as light as possible, and if pressurized, use a lower tank and engine pressure than otherwise possible. ML 3rd stage will have an engine pressure of 2-3 atm and 2nd stage about 5 atm. OK in vacuum.
With a 1st stage velocity low enough recovery at the launch site is fairly easy, making that highly re-usable whether the upper stages are.


You can lift more payload using two stages than with a single stage, but that doesn't mean that a SSTO is not possible, any more than the fact that you can lift more payload with three stages than two means TSTO is not possible.
Note also that even though you can lift more payload with three stages than with two that doesn't mean that two stages can't have operational and cost advantages over a three stage design, as is evidenced by the SpaceX design for the Falcon 9.
My point is the same is true for SSTO compared to TSTO. SSTO's can have operational and cost advantages over TSTO's.
In regards to the air drag such a SSTO would have to withstand, according to this news release Boeing has done wind tunnel tests on their proposed X-37B derived 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 ... emid=30110


Since the wind tunnel tests have been completed I interpreted this to mean the air drag issue has already been taken into account.


Bob Clark

_________________
Single-stage-to-orbit was already shown possible 50 years ago with the Titan II first stage.
Contrary to popular belief, SSTO's in fact are actually easy. Just use the most efficient engines
and stages at the same time, and the result will automatically be SSTO.
Blog: http://exoscientist.blogspot.com


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Post Re: An SSTO as "God and Robert Heinlein intended".   Posted on: Sun Jul 10, 2011 4:17 pm
marsmission2020 wrote:
Did you even review the structural innovations on the "create the future"
Website???
Search innovation on the create the future contest site
MSME 1985
Stress and fatigue designer on the x-33 and magnum launch vehicle projects
747 f/a-18 and other airframes



I'd like to read that. Could you provide the link?


Bob Clark

_________________
Single-stage-to-orbit was already shown possible 50 years ago with the Titan II first stage.
Contrary to popular belief, SSTO's in fact are actually easy. Just use the most efficient engines
and stages at the same time, and the result will automatically be SSTO.
Blog: http://exoscientist.blogspot.com


Last edited by RGClark on Sun Jul 10, 2011 4:28 pm, edited 1 time in total.



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Post Re: An SSTO as "God and Robert Heinlein intended".   Posted on: Sun Jul 10, 2011 4:26 pm
ckpooley wrote:
You missed the 2 points made in my previous post. They are not trivial.
They completely rule out SSTO as an actuality.


No they don't. There is a reason why the rocket equation is so widely used to estimate the capability of a launch system - because it works. Note most importantly that it takes into account air drag loss. Remember the original Atlas used balloon tank design to have a very lightweight designed stage that reached all the way to orbit. Two booster engines had to be dropped off because the engines were of low efficiency then, not because the structure could not withstand the aerodynamic stress.
With the more efficient engines available now, we would have both a lighter engine weight AND a higher Isp, resulting in a orbit capable single stage.


Bob Clark

_________________
Single-stage-to-orbit was already shown possible 50 years ago with the Titan II first stage.
Contrary to popular belief, SSTO's in fact are actually easy. Just use the most efficient engines
and stages at the same time, and the result will automatically be SSTO.
Blog: http://exoscientist.blogspot.com


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Post Re: An SSTO as "God and Robert Heinlein intended".   Posted on: Tue Jul 12, 2011 5:50 am
RGClark wrote:
...this was using the low efficiency engines available in the early 60's. Let's swap these out for the high efficiency NK-33 [1]. The sustainer engine used was the LR89-5 [2] at 720 kg. At 1,220 kg the NK-33 weighs 500 kg more. So removing both the sustainer and booster engines to be replaced by the NK-33 our loaded mass becomes 117,526 kg and the dry mass 2,826 kg, and the mass ratio 41.6 (!).
For the trajectory-averaged Isp, notice this is not just the midpoint between the sea level and vacuum value, since most of the flight to orbit is at high altitude at near vacuum conditions. A problem with doing these payload to orbit estimates is the lack of a simple method for getting the average Isp over the flight for an engine, which inhibits people from doing the calculations to realize SSTO is possible and really isn't that hard. I'll use a guesstimate Ed Kyle uses, who is a frequent contributor to NasaSpaceFlight.com and the operator of the Spacelaunchreport.com site. Kyle takes the average Isp as lying 2/3rds of the way up from the sea level value to the vacuum value [3]. The sea level value of the Isp for the NK-33 is 297 s, and the vacuum value 331 s. Then from this guesstimate the average Isp is 297 + (2/3)(331 - 297) = 319.667, which I'll round to 320 s.
Using this average Isp and a 8,900 m/s delta-V for a flight to orbit, we can lift 4,200 kg to orbit:

320*9.8ln((117,526+4,200)/(2,826+4,200)) = 8,944 m/s.

This is a payload fraction of 3.5%, comparable to that of many multi-stage rockets.


Dr. John Schilling has a launch performance estimator on his company's web page based on a numerical formula:

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

There is a disclaimer on the page that for user-defined vehicles it is limited to only 3-stage vehicles, and indeed I found previously when I tried to use it on a SSTO it didn't supply an answer. However, recently I found it even gives an answer for an SSTO vehicle.

This is the answer I got when I used the numbers of the above example:

-------------------------------------------------------------
Mission Performance:
Launch Vehicle: User-Defined Launch Vehicle
Launch Site: Cape Canaveral / KSC
Destination Orbit: 200 x 200 km, 28 deg
Estimated Payload: 4319 kg
95% Confidence Interval: 3077 - 5820 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.

--------------------------------------------------------------


The estimator requires you to input an Isp and thrust for the engines. This is meant the vacuum Isp and thrust. The program takes into account the losses due to reduced exhaust velocity at sea level and low altitude.
For this case I used the 331 s vacuum Isp and 1,636 kN vacuum thrust of the NK-33.


Bob Clark

_________________
Single-stage-to-orbit was already shown possible 50 years ago with the Titan II first stage.
Contrary to popular belief, SSTO's in fact are actually easy. Just use the most efficient engines
and stages at the same time, and the result will automatically be SSTO.
Blog: http://exoscientist.blogspot.com


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Post Re: An SSTO as "God and Robert Heinlein intended".   Posted on: Sat Jul 16, 2011 10:34 pm
Correction. I've been informed by email that I exchanged the booster engine on the Atlas rocket I discuss below with the sustainer engine used on the main stage.
The Astronautix site is down now but this older version of Astronautix has the Atlas rocket:

SLV-3 Atlas / Agena B.
http://www.friends-partners.org/partner ... vgenab.htm

The sustainer engine used was the LR-105-5:

LR-105-5.
http://www.friends-partners.org/partner ... lr1055.htm

This engine is 260 kg lighter than the booster engine so from the calculations of the payload to orbit subtract off 260 kg.
The payload possible is still several thousand kg and it is still the case that the payload to dry mass ratio is greater than one, which is better than any rocket that has ever reached orbit.


Bob Clark


RGClark wrote:
RGClark wrote:
SSTO is not a bad word. It is a very good word. It is my contention that the reason why launch costs are so high, the reason why we don't have passenger access to space as routine as say trans-Pacific flights is that the idea has been promulgated that SSTO is impossible. That is not the case. In fact it is easy, IF you do it in the right way. The right way is summarized in that one simple sentence at the end of my sig file.
We all know that to get a good payload to space you want a high efficiency engine. And we all know we want to use lightweight structures so the weight savings can go to increased payload. So you would think it would be obvious to use both these ideas to maximize the payload to orbit, right?
And indeed both have been used together - for upper stages. Yet this fundamentally obvious concept still has not been used for first stages. It is my thesis that if you do this then what you wind up with will automatically be SSTO capable. This is true for either kerosene fueled or hydrogen fueled stages.
Part of the misinformation that has been promulgated is that mass ratio for SSTO's is some impossible number. This is false. We've had rocket stages with the required mass ratio's since the 60's, nearly 50 years, both for kerosene and hydrogen fueled. Another part of the misinformation is that it would require some unknown high energy fuel and engine to accomplish. This is false. The required engines have existed since the 70's, nearly 40 years, both for kerosene and hydrogen fueled.
What has NOT been done is to marry the two concepts together for first stages. All you need to do is swap out the low efficiency engines that have been used for the high mass ratio stages and replace them with the high efficiency engines. It really is that simple.
This makes possible small, low cost orbital vehicles that could transport the same number of passengers as the space shuttle, about 7, but would have a comparable cost to a mid-sized business jet, a few tens of millions of dollars.
Then once you have the SSTO's they make your staged vehicles even better because you can carry greater payload when they are used for the individual stages of the multi-staged vehicle.


In disseminating the false dogma that SSTO's are not possible it is sometimes said instead that they are not practical because the payload fraction is so small. Even this is false. And indeed this is just as damaging as making the false statement they are not possible because the statements are often conflated into meaning the same thing. So when those in the industry make the statement they are not "practical", meaning actually they are doable but not economical, this becomes interpreted among many space enthusiasts and even many in the industry as meaning it would require some revolutionary advance to make them possible.
The fact that you can carry significant payload to orbit using SSTO's can be easily confirmed by anyone familiar with the rocket equation. To get a SSTO with significant payload using efficient kerosene engines you need a mass ratio of about 20 to 1. And to get a SSTO with significant payload using efficient hydrogen engines you need a mass ratio of about 10 to 1. Both of these the high mass ratio stages and the high efficiency engines for both kerosene and hydrogen have existed for decades now.
See this list of rocket stages:

Stages Index.
http://www.astronautix.com/stages/index.htm

Among the kerosene-fueled stages you see that several among the Atlas and Delta family have the required mass ratio. However, for the early Atlas stages you have to be aware of the type of staging system they used. They had drop-off booster engines and a main central engine, called the sustainer that continued all the way to orbit. But even when you take this into account you see these highly weight optimized stages had surprisingly high mass ratios.
See for instance the Atlas Agena SLV-3:

Atlas Agena SLV-3 Lox/Kerosene propellant rocket stage. Loaded/empty mass 117,026/2,326 kg. Thrust 386.30 kN. Vacuum specific impulse 316 seconds.
Cost $ : 14.500 million. Semistage: LR89-5. Semistage Thrust (vac): 1,644.960 kN (369,802 lbf). Semistage Thrust (vac): 167,740 kgf. Semistage specific impulse: 290 sec. Semistage Burn time: 120 sec. Semistage specific impulse (sl): 256 sec. Semistage Jettisonable Mass: 3,174 kg (6,997 lb). Semistage- number engines: 2. Semistage: Atlas MA-3.

Status: Out of production.
Gross mass: 117,026 kg (257,998 lb).
Unfuelled mass: 2,326 kg (5,127 lb).
Height: 20.67 m (67.81 ft).
Diameter: 3.05 m (10.00 ft).
Span: 4.90 m (16.00 ft).
Thrust: 386.30 kN (86,844 lbf).
Specific impulse: 316 s.
Specific impulse sea level: 220 s.
Burn time: 265 s.
Number: 140 .

http://www.astronautix.com/stages/atlaslv3.htm

Looking at only the loaded/empty mass you would think this stage had a mass ratio close to 50 to 1. But that is only including the sustainer engine. The more relevant ratio would be when you add in the mass of the booster engines to the dry mass since they are required to lift the vehicle off the pad. These are listed as the jettisonable mass at 3,174 kg. This makes the loaded mass now 117,026 + 3,174 = 120,200 and the dry mass 2,326 + 3,174 = 5,500 kg, for a mass ratio of 21.85.
But this was using the low efficiency engines available in the early 60's. Let's swap these out for the high efficiency NK-33 [1]. The sustainer engine used was the LR89-5 [2] at 720 kg. At 1,220 kg the NK-33 weighs 500 kg more. So removing both the sustainer and booster engines to be replaced by the NK-33 our loaded mass becomes 117,526 kg and the dry mass 2,826 kg, and the mass ratio 41.6 (!).
For the trajectory-averaged Isp, notice this is not just the midpoint between the sea level and vacuum value, since most of the flight to orbit is at high altitude at near vacuum conditions. A problem with doing these payload to orbit estimates is the lack of a simple method for getting the average Isp over the flight for an engine, which inhibits people from doing the calculations to realize SSTO is possible and really isn't that hard. I'll use a guesstimate Ed Kyle uses, who is a frequent contributor to NasaSpaceFlight.com and the operator of the Spacelaunchreport.com site. Kyle takes the average Isp as lying 2/3rds of the way up from the sea level value to the vacuum value [3]. The sea level value of the Isp for the NK-33 is 297 s, and the vacuum value 331 s. Then from this guesstimate the average Isp is 297 + (2/3)(331 - 297) = 319.667, which I'll round to 320 s.
Using this average Isp and a 8,900 m/s delta-V for a flight to orbit, we can lift 4,200 kg to orbit:

320*9.8ln((117,526+4,200)/(2,826+4,200)) = 8,944 m/s.

This is a payload fraction of 3.5%, comparable to that of many multi-stage rockets.
Note in fact that this has a very good value for a ratio that I believe should be regarded as a better measure, i.e., figure of merit, for the efficiency of a orbital vehicle. This is the ratio of the payload to the total dry mass of the vehicle. The reason why this is a good measure is because actually the cost of the propellant is a minor component for the cost of an orbital rocket. The cost is more accurately tracked by the dry mass and the vehicle complexity. Note that SSTO's in not having the complexity of staging are also good on the complexity scale.
For the ratio of the payload to dry mass you see this is greater than 1 for this SSTO. This is important because for every orbital vehicle I looked at, and possibly for every one that has existed, this ratio is going in the other direction: the vehicle dry mass is greater than the payload carried. Often it is much greater. For instance for the space shuttle system, the vehicle dry mass is more than 12 times that of the payload.
This good payload fraction and even better payload to dry mass ratio was just by using the engine in its standard configuration, no altitude compensation. However, for a SSTO you definitely would want to use altitude compensation. Dr. Bruce Dunn in his report "Alternate Propellants for SSTO Launchers" [4] estimates an average Isp of 338.3 s for high performance kerosene engines when using altitude compensation. Then we could lift 5,500 kg to orbit:

338.3*9.8ln((117,526+5,500)/(2,826+5,500)) = 8,928 m/s.

But kerosene is not the most energetic hydrocarbon fuel you could use. Dunn in his report estimates an average Isp of 352 s for methylacetyene using altitude compensation. This would allow a payload of 6,500 kg : 352*9.8ln((117,526+6,500)/(2,826+6,500)) = 8,926 m/s.


Bob Clark


REFERENCES.

1.)NK-33.
http://www.astronautix.com/engines/nk33.htm

2.)LR89-5.
http://www.astronautix.com/engines/lr895.htm

3.)Re: EELV Solutions for VSE.
Reply #269 on: 11/05/2007 09:20 PM
http://forum.nasaspaceflight.com/index. ... #msg208875

4.)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

Disclaimer: the citing of a particular reference should not be construed as an endorsement by the cited authors of the viewpoint expressed herein.

_________________
Single-stage-to-orbit was already shown possible 50 years ago with the Titan II first stage.
Contrary to popular belief, SSTO's in fact are actually easy. Just use the most efficient engines
and stages at the same time, and the result will automatically be SSTO.
Blog: http://exoscientist.blogspot.com


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Post Re: SSTO would have made possible A. C. Clarke's vision of 2001.   Posted on: Tue Jul 19, 2011 7:32 pm
This discussion thread on the SecretProjects forum, showed such
SSTO's were already being proposed in the 60's, as well as ambitious
lunar exploration proposals as exemplified by the lunar bases in the
film, 2001:

ROMBUS, Pegasus, Ithacus .
http://www.secretprojects.co.uk/forum/i ... pic=4577.0

We didn't have the required high efficiency kerosene or hydrogen
engines in the 60's. But we did in the 70's with the NK-33 for
kerosene and the SSME's for hydrogen.


Bob Clark




RGClark wrote:
Space Travel: The Path to Human Immortality?
Space exploration might just be the key to human beings surviving mass genocide, ecocide or omnicide.
July 24, 2009
On December 31st, 1999, National Public Radio interviewed the futurist and science fiction genius Arthur C. Clarke. Since the author had forecast so many of the 20th Century's most fundamental developments, the NPR correspondent asked Clarke if anything had happened in the preceding 100 years that he never could have anticipated. "Yes, absolutely," Clarke replied, without a moment's hesitation. "The one thing I never would have expected is that, after centuries of wonder and imagination and aspiration, we would have gone to the moon ... and then stopped."
http://www.alternet.org/news/141518/spa ... mortality/

I remember thinking when I first saw 2001 as a teenager and could appreciate it more, I thought it was way too optimistic. We could never have huge rotating space stations and passenger flights to orbit and Moon bases and nuclear-powered interplanetary ships by then.
That's what I thought and probably most people familiar with the space program thought that. And I think I recall Clarke saying once that the year 2001 was selected as more a rhetorical, artistic flourish rather than being a prediction, 2001 being the year of the turn of the millennium (no, it was NOT in the year 2000.)
However, I've now come to the conclusion those could indeed have been possible by 2001. I don't mean the alien monolith or the intelligent computer, but the spaceflights shown in the film.
It all comes down to SSTO's. As I argued above these could have led and WILL lead to the price to orbit coming down to the $100 per kilo range. The required lightweight stages existed since the 60's and 70's for kerosene with the Atlas and Delta stages, and for hydrogen with the Saturn V upper stages. And the high efficiency engines from sea level to vacuum have existed since the 70's with the NK-33 for kerosene, and with the SSME for hydrogen.
The kerosene SSTO's could be smaller and cheaper and would make possible small orbital craft in the price range of business jets, at a few tens of millions of dollars. These would be able to carry a few number of passengers/crew, say of the size of the Dragon capsule. But in analogy with history of aircraft these would soon be followed by large passenger craft.
However, the NK-33 was of Russian design, while the required lightweight stages were of American design. But the 70's was the time of detente, with the Apollo-Soyuz mission. With both sides realizing that collaboration would lead to routine passenger spaceflight, it is conceivable that they could have come together to make possible commercial spaceflight.
There is also the fact that for the hydrogen fueled SSTO's, the Americans had both the required lightweight stages and high efficiency engines, though these SSTO's would have been larger and more expensive. So it would have been advantageous for the Russians to share their engine if the American's shared their lightweight stages.
For the space station, many have soured on the idea because of the ISS with the huge cost overruns. But Bigelow is planning on "space hotels" derived from NASA's Transhab concept. These provide large living space at lightweight. At $100 per kilo launch costs we could form large space stations from the Transhabs linked together in modular fashion, financed purely from the tourism interests. Remember the low price to orbit allows many average citizens to pay for the cost to LEO.
The Transhab was developed in the late 90's so it might be questionable that the space station could be built from them by 2001. But remember in the film the space station was in the process of being built. Also, with large numbers of passengers traveling to space it seems likely that inflatable modules would have been thought of earlier to house the large number of tourists who might want a longer stay.
For the extensive Moon base, judging from the Apollo missions it might be thought any flight to the Moon would be hugely expensive. However, Robert Heinlein once said: once you get to LEO you're half way to anywhere in the Solar System. This is due to the delta-V requirements for getting out of the Earth's gravitational well compared to reaching escape velocity.
It is important to note then SSTO's have the capability once refueled in orbit to travel to the Moon, land, and return to Earth on that one fuel load. Because of this there would be a large market for passenger service to the Moon as well. So there would be a commercial justification for Bigelow's Transhab motels to also be transported to the Moon.
Initially the propellant for the fuel depots would have to be lofted from Earth. But we recently found there was water in the permanently shadowed craters on the Moon. Use of this for propellant would reduce the cost to make the flights from LEO to the Moon since the delta-V needed to bring the propellant to LEO from the lunar surface is so much less than that needed to bring it from the Earth's surface to LEO.
This lunar derived propellant could also be placed in depots in lunar orbit and at the Lagrange points. This would make easier flights to the asteroids and the planets. The flights to the asteroids would be especially important for commercial purposes because it is estimated even a small sized asteroid could have trillions of dollars worth of valuable minerals. The availability of such resources would make it financially profitable to develop large bases on the Moon for the sake of the propellant.
Another possible resource was recently discovered on the Moon: uranium. Though further analysis showed the surface abundance to be much less than in Earth mines, it may be that there are localized concentrations just as there are on Earth. Indeed this appears to be the case with some heavy metals such as silver and possibly gold that appear to be concentrated in some polar craters on the Moon.
So even if the uranium is not as abundant as in Earth mines, it may be sufficient to be used for nuclear-powered spacecraft. Then we wouldn't have the problem of large amounts of nuclear material being lofted on rockets on Earth. The physics and engineering of nuclear powered rockets have been understood since the 60's. The main impediment has been the opposition to launching large amounts of radioactive material from Earth into orbit above Earth. Then we very well could have had nuclear-powered spacecraft launching from the Moon for interplanetary missions, especially when you consider the financial incentive provided by minerals in the asteroids of the asteroid belt.


Bob Clark

_________________
Single-stage-to-orbit was already shown possible 50 years ago with the Titan II first stage.
Contrary to popular belief, SSTO's in fact are actually easy. Just use the most efficient engines
and stages at the same time, and the result will automatically be SSTO.
Blog: http://exoscientist.blogspot.com


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Post Re: An SSTO as "God and Robert Heinlein intended".   Posted on: Tue Jul 19, 2011 7:55 pm
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.


Bob Clark

_________________
Single-stage-to-orbit was already shown possible 50 years ago with the Titan II first stage.
Contrary to popular belief, SSTO's in fact are actually easy. Just use the most efficient engines
and stages at the same time, and the result will automatically be SSTO.
Blog: http://exoscientist.blogspot.com


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Post Re: An SSTO as "God and Robert Heinlein intended".   Posted on: Sat Jul 23, 2011 4:02 am
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 the larger versions become fully orbital vehicles.

I discuss this in regard to the Air Forces's X-37B and Sierra Nevada's Dream Chaser here:

Newsgroups: sci.space.policy, sci.astro, sci.physics, sci.space.history
From: Robert Clark <rgregorycl...@yahoo.com>
Date: Fri, 22 Jul 2011 15:09:14 -0700 (PDT)
Subject: Re: A kerosene-fueled X-33 as a single stage to orbit vehicle.
http://groups.google.com/group/sci.spac ... f1ea?hl=en

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 liquid fueled engines produces a SSTO.


Bob Clark

_________________
Single-stage-to-orbit was already shown possible 50 years ago with the Titan II first stage.
Contrary to popular belief, SSTO's in fact are actually easy. Just use the most efficient engines
and stages at the same time, and the result will automatically be SSTO.
Blog: http://exoscientist.blogspot.com


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Post Re: An SSTO as "God and Robert Heinlein intended".   Posted on: Sat Jul 23, 2011 7:13 pm
RGClark wrote:
Then swapping out the hybrid engines of SpaceShipTwo for high efficiency liquid fueled engines produces a SSTO.

No it doesn't, not even close. SS2 is a long way away from the MR needed for an SSTO.

If you really think it is, demonstrate it with math, rather than hand waving.


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Post Re: SSTO would have made possible A. C. Clarke's vision of 2001.   Posted on: Fri Aug 05, 2011 11:19 am
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/lookin ... aking.html

August 02, 2011
Elon Musk of Spacex talks about a Reusable Falcon Heavy to get to $50 a pound to space.
Quote:
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.
{emphasis added - B.C.}
http://nextbigfuture.com/2011/08/elon-m ... about.html


Bob Clark

RGClark wrote:
Space Travel: The Path to Human Immortality?
Space exploration might just be the key to human beings surviving mass genocide, ecocide or omnicide.
July 24, 2009
On December 31st, 1999, National Public Radio interviewed the futurist and science fiction genius Arthur C. Clarke. Since the author had forecast so many of the 20th Century's most fundamental developments, the NPR correspondent asked Clarke if anything had happened in the preceding 100 years that he never could have anticipated. "Yes, absolutely," Clarke replied, without a moment's hesitation. "The one thing I never would have expected is that, after centuries of wonder and imagination and aspiration, we would have gone to the moon ... and then stopped."
http://www.alternet.org/news/141518/spa ... mortality/

I remember thinking when I first saw 2001 as a teenager and could appreciate it more, I thought it was way too optimistic. We could never have huge rotating space stations and passenger flights to orbit and Moon bases and nuclear-powered interplanetary ships by then.
That's what I thought and probably most people familiar with the space program thought that. And I think I recall Clarke saying once that the year 2001 was selected as more a rhetorical, artistic flourish rather than being a prediction, 2001 being the year of the turn of the millennium (no, it was NOT in the year 2000.)
However, I've now come to the conclusion those could indeed have been possible by 2001. I don't mean the alien monolith or the intelligent computer, but the spaceflights shown in the film.
It all comes down to SSTO's. As I argued above these could have led and WILL lead to the price to orbit coming down to the $100 per kilo range. The required lightweight stages existed since the 60's and 70's for kerosene with the Atlas and Delta stages, and for hydrogen with the Saturn V upper stages. And the high efficiency engines from sea level to vacuum have existed since the 70's with the NK-33 for kerosene, and with the SSME for hydrogen.
The kerosene SSTO's could be smaller and cheaper and would make possible small orbital craft in the price range of business jets, at a few tens of millions of dollars. These would be able to carry a few number of passengers/crew, say of the size of the Dragon capsule. But in analogy with history of aircraft these would soon be followed by large passenger craft.
However, the NK-33 was of Russian design, while the required lightweight stages were of American design. But the 70's was the time of detente, with the Apollo-Soyuz mission. With both sides realizing that collaboration would lead to routine passenger spaceflight, it is conceivable that they could have come together to make possible commercial spaceflight.
There is also the fact that for the hydrogen fueled SSTO's, the Americans had both the required lightweight stages and high efficiency engines, though these SSTO's would have been larger and more expensive. So it would have been advantageous for the Russians to share their engine if the American's shared their lightweight stages.
For the space station, many have soured on the idea because of the ISS with the huge cost overruns. But Bigelow is planning on "space hotels" derived from NASA's Transhab concept. These provide large living space at lightweight. At $100 per kilo launch costs we could form large space stations from the Transhabs linked together in modular fashion, financed purely from the tourism interests. Remember the low price to orbit allows many average citizens to pay for the cost to LEO.
The Transhab was developed in the late 90's so it might be questionable that the space station could be built from them by 2001. But remember in the film the space station was in the process of being built. Also, with large numbers of passengers traveling to space it seems likely that inflatable modules would have been thought of earlier to house the large number of tourists who might want a longer stay.
For the extensive Moon base, judging from the Apollo missions it might be thought any flight to the Moon would be hugely expensive. However, Robert Heinlein once said: once you get to LEO you're half way to anywhere in the Solar System. This is due to the delta-V requirements for getting out of the Earth's gravitational well compared to reaching escape velocity.
It is important to note then SSTO's have the capability once refueled in orbit to travel to the Moon, land, and return to Earth on that one fuel load. Because of this there would be a large market for passenger service to the Moon as well. So there would be a commercial justification for Bigelow's Transhab motels to also be transported to the Moon.
Initially the propellant for the fuel depots would have to be lofted from Earth. But we recently found there was water in the permanently shadowed craters on the Moon. Use of this for propellant would reduce the cost to make the flights from LEO to the Moon since the delta-V needed to bring the propellant to LEO from the lunar surface is so much less than that needed to bring it from the Earth's surface to LEO.
This lunar derived propellant could also be placed in depots in lunar orbit and at the Lagrange points. This would make easier flights to the asteroids and the planets. The flights to the asteroids would be especially important for commercial purposes because it is estimated even a small sized asteroid could have trillions of dollars worth of valuable minerals. The availability of such resources would make it financially profitable to develop large bases on the Moon for the sake of the propellant.
Another possible resource was recently discovered on the Moon: uranium. Though further analysis showed the surface abundance to be much less than in Earth mines, it may be that there are localized concentrations just as there are on Earth. Indeed this appears to be the case with some heavy metals such as silver and possibly gold that appear to be concentrated in some polar craters on the Moon.
So even if the uranium is not as abundant as in Earth mines, it may be sufficient to be used for nuclear-powered spacecraft. Then we wouldn't have the problem of large amounts of nuclear material being lofted on rockets on Earth. The physics and engineering of nuclear powered rockets have been understood since the 60's. The main impediment has been the opposition to launching large amounts of radioactive material from Earth into orbit above Earth. Then we very well could have had nuclear-powered spacecraft launching from the Moon for interplanetary missions, especially when you consider the financial incentive provided by minerals in the asteroids of the asteroid belt.


Bob Clark

_________________
Single-stage-to-orbit was already shown possible 50 years ago with the Titan II first stage.
Contrary to popular belief, SSTO's in fact are actually easy. Just use the most efficient engines
and stages at the same time, and the result will automatically be SSTO.
Blog: http://exoscientist.blogspot.com


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Post Re: An SSTO as "God and Robert Heinlein intended".   Posted on: Thu Aug 11, 2011 2:59 pm
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 jets, 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/Pra ... e_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.spac ... 5a22?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 ... -and/15354

13.)Reusable Launch System.
http://en.wikipedia.org/wiki/Reusable_l ... al_landing

14.)Landing gear weight (Gary Hudson; George Herbert; Henry Spencer).
http://yarchive.net/space/launchers/lan ... eight.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/ ... 1-4619.pdf

_________________
Single-stage-to-orbit was already shown possible 50 years ago with the Titan II first stage.
Contrary to popular belief, SSTO's in fact are actually easy. Just use the most efficient engines
and stages at the same time, and the result will automatically be SSTO.
Blog: http://exoscientist.blogspot.com


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Post Re: An SSTO as "God and Robert Heinlein intended".   Posted on: Fri Aug 26, 2011 8:30 am
RGClark wrote:
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:

viewtopic.php?p=45803#p45803

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

_________________
Single-stage-to-orbit was already shown possible 50 years ago with the Titan II first stage.
Contrary to popular belief, SSTO's in fact are actually easy. Just use the most efficient engines
and stages at the same time, and the result will automatically be SSTO.
Blog: http://exoscientist.blogspot.com


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