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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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Post Re: An SSTO as "God and Robert Heinlein intended".   Posted on: Fri Aug 26, 2011 8:36 am
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

_________________
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 Sep 13, 2011 3:57 pm
I saw this discussed on a space oriented forum:

WSJ: Europe Ends Independent Pursuit of Manned Space Travel.
"LE BOURGET, France—Europe appears to have abandoned all hope of
independently pursuing human space exploration, even as the region's
politicians and aerospace industry leaders complain about shrinking
U.S. commitment to various space ventures.
"After years of sitting on the fence regarding a separate, pan-
European manned space program, comments by senior government and
industry officials at the Paris Air Show here underscore that budget
pressures and other shifting priorities have effectively killed that
longtime dream."
http://www.orbiter-forum.com/showthread.php?t=23006

In this post I discussed getting a SSTO by replacing the Vulcain
engine on the Ariane 5 core with a SSME:

Newsgroups: sci.space.policy, sci.astro, sci.physics,
sci.space.history
From: Robert Clark <rgregorycl...@yahoo.com>
Date: Wed, 23 Feb 2011 10:14:42 -0800 (PST)
Subject: Re: Some proposals for low cost heavy lift launchers.
http://groups.google.com/group/sci.phys ... 269f?hl=en

However, in point of fact Europe can produce a manned launch vehicle
from currently *existing*, European components. This will consist of
the Ariane 5 and three Vulcain engines. The calculations below use the
Ariane 5 generic "G" version. You might need to add another Vulcain
for the larger evolution "E" version of the Ariane 5 core.
In a following post I'll also show that the Hermes spaceplane also
can become a SSTO by filling the entire fuselage aft of the cockpit
with hydrocarbon propellant.
The impetus for trying the calculation for a Ariane 5 core based SSTO
using Vulcains instead of the SSME was from a report by SpaceX that
you could get the same performance from a planned heavy lift first
stage using a lower performance Merlin 2 compared to the high
performance RS-84 engine. The reason was the lower Isp of the Merlin
was made up for by its lower weight.

THIS IS A VERY IMPORTANT FACT BECAUSE WHAT IT MEANS IS THAT YOU DON'T
NEED THE HIGH PERFORMANCE ENGINES TO GET THE SSTO. YOU CAN USE ENGINES
OF LOWER CHAMBER PRESSURE AND SIMPLER COMBUSTION CYCLES, SUCH AS THE
VULCAIN WITH A CA. 100 BAR COMBUSTION PRESSURE AND A GAS GENERATOR
CYCLE. THIS MEANS THE ENGINES ARE CHEAPER, EASIER TO MAKE REUSABLE,
REQUIRE LESS ROUTINE MAINTENANCE, AND CAN LAST FOR MANY RESTARTS.

In the discussion of the Ariane/Vulcain SSTO below, I note you can
get a prototype, test vehicle quite quickly since the components are
already existing. To improve the payload though you would want to use
altitude compensation on the Vulcains. In a following post I'll
discuss some methods of altitude compensation.
In regards to achieving this at low cost, I think the most important
accomplishment of SpaceX might turn out to be that they showed in
stark terms that privately financed spacecraft, both launchers and
crew capsules, can be accomplished at 1/10th the developmental cost of
government financed ones. Imagine a manned, reusable orbital launcher,
for example, instead of costing, say, $3 billion, only costing $300
million to develop.
Here's how you can get an all European manned SSTO using the Ariane 5
core stage but with Vulcain engines this time. Note that this is one
that can be produced from currently existing components, aside from
the capsule, so at least an unmanned prototype vehicle can be
manufactured and tested in the short term and at lowered development
cost.
We'll use three Vulcain 2's instead of the 1 normally used with the
Ariane 5 core stage. There are varying specifications given on the
Vulcain 2 depending on the source. I'll use the Astronautix site:

Vulcain 2.
http://www.astronautix.com/engines/vulcain2.htm

From the sea level thrust given there, using three Vulcain 2's will
give us one engine out capability. The weight is given as 1,800 kg. So
adding on two will take the dry mass from 12 mT to 15.6 mT.
To calculate the delta-V achieved I'll use the idea again to just use
the vacuum Isp, but adding the loss due to back pressure onto the
delta-V required for orbit, as I discussed previously. However, here
for hydrogen fuel which has higher gravity loss, I'll use a higher
required delta-V of 9,400 m/s when you add on the back pressure loss.
With the vacuum Isp given for the Vulcain 2 of 434 s, we get a payload
of 3.8 mT:

434*9.8ln(1+158/(15.6 + 3.8 )) = 9,412 m/s.

Note this is just using the standard nozzle Isp for the Vulcain, no
altitude compensation. So this could be tested, like, tomorrow.
However, for a SSTO you definitely want to use altitude compensation.
Using engine performance programs such as ProPEP we can calculate that
using long nozzles, you can get a vacuum Isp of 470 s for this engine.
As a point of comparison of how high an Isp you can get even with a
low chamber pressure engine as long as you have a long nozzle, or
equivalent, note that the RL10-B2 with a ca. 250 to 1 area ratio, and
only a ca. 40 bar chamber pressure, gets a 465 s vacuum Isp. So we'll
assume we can get a comparable Isp by using altitude compensation.
This allows us to get payload of 8 mT:

470*9.8ln(1+158/(15.6+8) = 9,400 m/s.

This allows us to add a Dragon-sized capsule and also the reentry and
landing systems to make it reusable.

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 Sep 17, 2011 5:36 pm
In a post from Aug. 11th I argued that small, low cost SSTO's are doable
now using lightweight design and high efficiency engines. However,
I was surprised to find after doing the calculation you don't even need
the high efficiency engines to get the SSTO. The low efficiency SpaceX
Merlin engines would be sufficient for example, IF you have altitude compensation.
The impetus for trying the calculation was from a report by SpaceX
that you could get the same performance from a planned heavy lift
first stage using a lower performance Merlin 2 engine compared to the
high performance RS-84 engine. The reason was the lower Isp of the
Merlin was made up for by its lower weight:

SpaceX Propulsion.
http://images.spaceref.com/news/2010/Sp ... ulsion.pdf

Now note that the biggest single contributor to the vacuum Isp of an
engine is not the chamber pressure, but the nozzle length. For
example, the Merlin Vacuum raises its vacuum Isp to 342 s from the 304
s Isp of the Merlin 1C by having a longer nozzle, even though the
chamber pressure remains the same, ca. 100 bar.
So I'll redo the calculation for the SSTO using the SpaceX Falcon 1
first stage but using Merlin engines this time. We'll assume that
using altitude compensation we are able to get an engine with the same
vacuum Isp as the Merlin Vacuum but able to launch from ground.
We'll use the soon to be introduced Merlin 1D:

SpaceX Plans To Be Top World Rocket Maker.
Aug 11, 2011
By Guy Norris
San Diego
http://www.aviationweek.com/aw/generic/story.jsp?id=news/awst/2011/08/08/AW_08_08_2011_p27-354586.xml&headline=SpaceX%20Plans%20To%20Be%20Top%20%20World%20Rocket%20Maker&channel=defense

Using the 160 to 1 thrust/weight ratio and 155,000 lbs. vacuum thrust
given, it has a mass of 970 lbs., 440 kg. However, this would make it
overpowered for the Falcon 1 first stage only. So we'll use two copies
of this stage powered by a single Merlin 1D.
The original Falcon 1 first stage with the Merlin 1C engine has a dry mass
of 1,360 kg. I estimated the mass of the Merlin 1C in the prior post to
be 650 kg. So without the engine, the stage weighs 710 kg. So two of
them will be 1,420 kg without engines, and adding on the Merlin 1D
engine gives this a mass of 1,860 kg.
The propellant mass of the two copies of the first stage is 43,080
kg. Then to calculate the payload that can be carried I'll again just
use the vacuum Isp and take the required delta-V as 9,150 m/s. We
conclude a payload of 1,140 kg can be lofted:

342*9.8ln(1 + 43,080/(1,860 + 1,140)) = 9,160.

Now we'll estimate how much the payload can be if we use a higher
energy density fuel such as methylacetylene and use lightweight
composites for the stage. I'll get a rough idea how high the Isp can
be for this case by assuming it is increased proportionally to the
same degree as for the high efficiency engine case. That is, using
methylacetylene in the high efficiency case resulted in increasing the
vacuum Isp to 384 s from the 360 s vacuum Isp for the kerosene.
Assuming the vacuum Isp will be increased to the same proportion here
gives us a vacuum Isp of 365 s for methylacetylene and the Merlin 1D
engine.
For the reduced stage weight using composites, assume again it will
be reduced by 40% aside from the engines. Then the stage weight with
the Merlin 1D engine will be .6*1,420 + 440 kg = 1,290 kg. Then will
be able to loft a payload of 2,320 kg:

365*9.8ln(1 + 43,080/(1,290 + 2,320)) = 9,160 m/s.

Also, quite likely SpaceX could make a half-size version of the
Merlin 1D engine. So you could use a single copy of the Falcon 1 first
stage. Then the payload would be approximately cut in half, 570 kg for
the kerosene/standard stage version and 1,160 kg for the
methylacetylene/composite stage version.

Note that low chamber pressure, low performance engines can also be
used to power the SSTO's is extremely important. Such engines have
less complicated combustion cycles and have to withstand much less
strenuous operating regimes. This makes them cheaper, simpler, easier
to maintain, and easier to make reusable. So the most costly component
of any rocket, the engines, become markedly cheaper for the proposed
SSTO.

What is key though is to come up with ways to get the needed altitude
compensation without adding on too much to the engine weight. In a
following post I'll discuss some methods this might be accomplished.


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: SSTO would have made possible A. C. Clarke's vision of 2001.   Posted on: Fri Dec 09, 2011 7:14 am
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.


Just saw this article on The Space Review discussing a recently discovered copy of a 1963 TV interview with Arthur C. Clarke:

The perils of spaceflight prediction.
by Jeff Foust
Monday, December 5, 2011
http://thespacereview.com/article/1981/1

In the interview Clarke gives some predictions of the future of space exploration. From the standpoint of the beginnings of human spaceflight, he suggests a manned Mars mission within 25 years, which would have been by 1988, and Moon bases by the end of the 20th century.
This turned out to be too optimistic. But as I argued above, this could indeed have been technically and even financially feasible: if it had been recognized that reusable SSTO's are possible and in fact aren't even really hard, we would have had routine, private spaceflight by the 1970s.
Such wide spread, frequent launches using reusable spacecraft would have cut the costs to space by two orders of magnitude, at least. This would then have made the costs of lunar bases and manned Mars missions well within the affordability range.
The important point is that the required high efficiency engines and lightweight stages for SSTO's already exist and have for decades. All that is required is to marry the two together. An expendable test SSTO could be produced, like, tomorrow. Just this one simple, cheap test would finally make clear the fact that routine spaceflight is already 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: Thu Jan 12, 2012 3:54 pm
Elon Musk has said he wants to cut the costs to space to the $100 to $200 per kg range by reusability. This is about a two order of magnitude reduction in cost. To put this in perspective, this is like a trans-atlantic flight that costs $1,000 suddenly being cut to cost $10 to $20.
Musk has said this transformation of the Falcon 9 to full reusability will be very hard. I don't believe it will be. But first, keep in mind how important that reduction in cost will be if it succeeds. If it succeeds then SpaceX will monopolize the launch business if the other launch companies do not field their own reusable vehicles. So there is a tremendous financial incentive for SpaceX to invest in reusability. Now, most in the industry believe reusability is very difficult for orbital vehicles and not even worth the expense. So if Musk reinforces that idea then he has a better chance at being able to field one without the other launch providers having one. And since they will not have even started to develop one, it will take them some time to catch up. The effect is that Musk will have a monopoly on all launches for at least a few years.
I don't know if that is Musk's intent in saying reusability is very hard. Actually I'm inclined to believe he is just saying what most in the industry believe including his own engineers. But a key reason why reusability is not very hard is because the cost in mass in reentry and landing systems is surprisingly low. In regards to the technical difficulty, there is none. We know how to do it as the shuttle orbiter and the X-37B and Dragon spacecraft has shown. I include the Dragon in the list of reusables because its heat shield showed minimal degradation on return. Musk has said the same heat shield could make hundreds of flights, at least to LEO.
I made an estimate before of about 28% of the landed mass has to go to reentry/landing systems. This was based on estimates of 15% for thermal protection, 10% for wings or for propellant for vertical landing, and 3% for landing gear. However, I said likely with modern materials this could be cut to half that. In fact, it might even be lower than 10%.

1.)Weight of thermal protection.

Robert Zubrin has given an estimate of 15% of the landed weight for the weight of thermal protection systems(TPS):

Reusable launch system.
http://en.wikipedia.org/wiki/Reusable_l ... at_shields

However, I gather this was in relation to the older capsules, Mercury, Gemini, Apollo, etc. Indeed the weight of the ablative heat shield on the Apollo capsule was about 15%:

Apollo Command/Service Module.
2.7 Specifications
http://en.wikipedia.org/wiki/Apollo_Com ... ifications

However, the space shuttle with its mostly silica tiles was able to reduce the TPS weight to about 8% of the maximum landing weight of 104,000 kg:

Space Shuttle thermal protection system.
3.3 Weight considerations.
http://en.wikipedia.org/wiki/Space_shut ... iderations

Also, for the X-37B the TUFROC leading edge material instead of the shuttles RCC and the TUFI AETB material instead of the shuttles silica tiles are either of equal or lower weight than the shuttles TPS materials while being tougher and requiring less maintenance:

X-37B Orbital Test Vehicle.
http://www.boeing.com/defense-space/ic/ ... b_otv.html

For ablative TPS, the PICA-X material used on the Dragon capsule weights about half the weight of the AVCOAT material used on the Apollo heat shield:

Re: Dragon v/s Orion.
http://forum.nasaspaceflight.com/index. ... #msg754168

while being able to still survive lunar and even Martian reentry speeds.

SpaceX has found that at least for LEO reentry speed judging from the minimal degradation on the Falcon 9/Dragon test flight, the PICA-X heat shield could be reused hundreds of times.

Also, for vertical powered landings a la the DC-X, you might not even need an extra heat shield for base first landings. One proposal for a VTVL SSTO uses low thrust during the descent as well as a high temperature-resistant aerospike nozzle to serve as the reentry thermal protection. You would need to retain more mass in propellant or some inert gas for this purpose though.
Another idea for a vertical landing vehicle would be to reenter head first. This was the preferred method of the Air Force since it provided increased cross-range. In that case you would have the blunt heat shield at the top of each stage. I thought this method would be unstable with the heavy engines now at the top during reentry, but since this was considered for the orbital version of the DC-X presumably this was solved.

2.)Weight of the wings and the landing gear.

For horizontal landing, a common estimate is that the weight of wings is 10% of the landed weight. This comes from aircraft examples though where the wings have to carry the weight of the fuel which can be as much as the dry weight of the aircraft itself or more.
An example where the propellant will not be carried in the wings and lightweight composites will be used is the Skylon. According to their released specifications the wing weight will be less than 2% of the take-off weight, which is the appropriate weight to compare to for a horizontal take-off vehicle:

The SKYLON Spaceplane.
by Richard Varvill and Alan Bond
Journal of the British Interplanetary Society, Vol. 57. pp. 22–32, 2004
p. 32.
http://www.reactionengines.co.uk/downlo ... _22-32.pdf

On that same page the landing gear weight is the only 1.5% of the take-off weight.
Then for a vertical take-off vehicle these low weight proportions should apply to the dry, landing weight.


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: Fri Jan 13, 2012 1:34 am
What about a mixed system? A device just like you are describing except initially launched via magnetic or electric linear acceleration? With space hubs you could refuel reload etc and taxi to the bottom of a slightly inclined very long runway that is the engine initially, initiating launch while at higher speeds would allow for much less fuel and reusability. The more structure on the ground the lower your weight and higher your payoff.

Is there any research into scramjets that run on a stable liquid fuel?

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Post Re: An SSTO as "God and Robert Heinlein intended".   Posted on: Fri Jan 13, 2012 10:38 am
The problem with getting lots of speed at low altitude (ie launch ramp) is the atmospheric friction is too high, and you either vapourise your craft or waste a lot of energy or don't gain much speed. Fine on the moon thoigh.

Lots of research on Scramjet tech, for example NASA, and an Australian university. Use Google.


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Post Re: An SSTO as "God and Robert Heinlein intended".   Posted on: Fri Jan 13, 2012 5:19 pm
What about having the launch site high in the mountains?

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Post Re: An SSTO as "God and Robert Heinlein intended".   Posted on: Fri Jan 13, 2012 11:29 pm
Nope. Won't make that much of a difference. While the drop in atmospheric pressure isn't linear, Even the top of Mt. Everest (~8km MSL) only reduces pressure by about half, its still significant. Your gun would have to be in the high stratosphere to note any serious reduction in drag.

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Post Re: An SSTO as "God and Robert Heinlein intended".   Posted on: Fri Jan 13, 2012 11:36 pm
That's not really true, everything is about margins. A rocket launched from the top of a mountain will have significantly higher performance than one launched at sea level; the gain due to reduced drag losses is part, but there are also reduced gravity losses and a significant impact on the Isp of the engine and how far the first stage can be expanded.

For a suborbital trajectory, increasing the launch altitude by a thousand feet can increase the apogee by 3000 feet.


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Post Re: An SSTO as "God and Robert Heinlein intended".   Posted on: Sat Jan 14, 2012 5:19 am
For sub-orbital its a percentage point or two. For orbital its a drop in the bucket. And a vertical rocket's assent profile means its above most of the drag inducing atmosphere before it gets hypersonic anyway.

But for the context at hand, horizontally launching something with its max velocity, its a big problem that even being at the top of Mt. Everest won't solve (my point). Might be useful for getting something started withing the operating range of a RAM or SCRAM jet etc., but their are easier and simpler ways of peeling that banana.


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Post Re: An SSTO as "God and Robert Heinlein intended".   Posted on: Tue Jan 24, 2012 5:35 pm
JamesHughes wrote:
Why use SSTO orbit? Why not launch using two stages, the first dropping off and being recovered (vertical landing?) Then the second stage is in effect a DC-X style craft, still using vertical descent, but because of the greatly reduced mass, requiring less fuel, so more cargo.
In a word, efficiency. 2 craft weigh more than one, and carry less than twice the fuel, because of the square-cube law. If you have 2 square card tables, you can put 4 chairs around them, but if you scoot them together, you have room for 6 place settings. Now, get 4 tables, which have 16 spacees around them seperately, or arrange them in a square, and there are only 8.

The chairs are the material you inevitably have to wrap around your fuel tank, habitat volume, and associated gear. Now, multiply that analogy to 3 dimensions, and you see why it's called the square-cube law. If you double the diameter, you square (x4) the surface area, and cube (x8) the internal volume. If you double the tank, you double the weight, and the capacity linearly, instead of exponentially. Now apply that to drag, mainenance, and all the other associated complexities than make Rocket Science a cliche' for mind boggling difficulty, and you understand why an SSTO (Or OStEO, for One Stage to Earth Orbit) is something you'd want.

Only by exponential scaling can you acceive the lift/mass ratios neccessary to make spaceflight reasonably affordable. That's why NACA, then Nasa originaly used big ass rockets with itty bitty payloads at the top. Multistaging was a fast, and dirty cheat with chemical rockets so they can lose mass as they climb. Also, we don't yet have a working engine that's efficient all the way up. A good rocket on the ground is not all that hot at high altitude, much less in the near vaccuum of the orbital environment.

Current dual stage designs, like most of the X-Prize participants, and the winner make up for this by using an air breathing stage in the atmosphere, and a rocket outside. This is a happy medium between a pure OStEO, and the immense rocket stacks of the early days. With wings, and jets, the booster can fly up to where a rocket is more efficient, but those wings, and extra engines are heavy, so you don't want to carry them all the way to orbit.

What I'd really like to try is piggy-backing an X-15 on an SR-71, but though off the shelf, it would be expensive. With modern upgrades (Not to mention adding an airlock), it could be a reasonably fast, and dirty personal shuttle to orbit.

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Post Re: An SSTO as "God and Robert Heinlein intended".   Posted on: Mon May 28, 2012 2:41 pm
New post to my blog:

The Coming SSTO's
http://exoscientist.blogspot.com/2012/0 ... sstos.html

Bob Clark

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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: Wed Jun 06, 2012 6:26 pm
I always liked that design, glad to see they're still working on it.

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Post Re: An SSTO as "God and Robert Heinlein intended".   Posted on: Sun Jul 22, 2012 12:41 pm
No, SSTO is not a four-letter word, though it is sometimes treated that way by those in the industry.
I've been arguing that SSTO's are actually easy because how to achieve them is perfectly obvious: use the most weight optimized stages and most Isp efficient engines at the same time, i.e., optimize both components of the rocket equation. But I've recently found it's even easier than that! It turns out you don't even need the engines to be of particularly high efficiency.
SpaceX is moving rapidly towards testing its Grasshopper scaled-down version of a reusable VTVL Falcon 9 first stage:

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

SpaceX will be duplicating in this what the DC-X accomplished in the early 90's. The DC-X was a scaled down, low altitude test vehicle for a full-scale SSTO VTVL vehicle. So could the full-sized Falcon 9 first stage act as a VTVL SSTO?
SpaceX deserves kudos for achieving a highly weight optimized Falcon 9 first stage at a 20 to 1 mass ratio. However, the Merlin 1C engine has an Isp no better than the engines we had in the early sixties at 304 s, and the Merlin 1D is only slightly better on the Isp scale at 310 s. This is well below the highest efficiency kerosene engines (Russian) we have now whose Isp's are in the 330's. So I thought that closed the door on the Falcon 9 first stage being SSTO.
However, I was surprised when I did the calculation that because of the Merlin 1D's lower weight the Falcon 9 first stage could indeed be SSTO. I'll use the Falcon 9 specifications estimated by GW Johnson, a former rocket engineer, now math professor:

WEDNESDAY, DECEMBER 14, 2011
Reusability in Launch Rockets.
http://exrocketman.blogspot.com/2011/12 ... ckets.html

The first stage propellant load is given as 553,000 lbs, 250,000 kg, and the dry weight as 30,000 lbs, 13,600 kg. The Merlin 1C mass hasn't been released, but I'll estimate it as 650 kg, from its reported thrust and thrust/weight ratio. The Merlin 1D mass has been estimated to be 450 kg. Then on replacing the 1C with the 1D we save 9*200 = 1,800kg from the dry weight to bring it to 11,800 kg.
The required delta v to orbit is frequently estimated as 30,000 feet per second for kerosene-fueled vehicles, 9,144 m/s. When calculating the delta v your rocket can achieve, you can just use your engines vacuum Isp since the loss of Isp at sea level is taken into account in the 30,000 fps number. Then this version of the Falcon 9 first stage could lift 1,200 kg to orbit:

310*9.81ln(1 + 250/(11.8 + 1.2)) = 9,145 m/s.

Then the Falcon 9 first stage could serve as a proof of principle SSTO on the switch to the Merlin 1D engine.


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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