SSTO's would have made possible Arthur C. Clarke's vision of 2001.

**Exoscientist** · 07-02-11, 02:33 PM

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/...n_immortality/

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 in the thread A kerosene-fueled X-33 as a single stage to orbit vehicle 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 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

**Exoscientist** · 07-19-11, 02:36 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/foru...p?topic=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

**Exoscientist** · 07-22-11, 10:50 PM

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.s...efcbf1ea?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

**billvon** · 07-23-11, 01:53 AM

Originally Posted by Exoscientist

Then swapping out the hybrid engines of SpaceShipTwo for high efficiency liquid fueled engines produces a SSTO.

Would take a lot more than a "high efficiency engine" to turn Spaceship Two into an SSTO. First you'd have to modify it so it could take off by itself; currently it needs a carrier ship to launch it. Next you'd have to greatly increase its propellant fraction. (Perhaps an external tank or two.) Then you'd have to come up with a thermal protection system capable of handling re-entry AND protecting those external tanks.

**Exoscientist** · 07-28-11, 10:17 PM

Originally Posted by Exoscientist

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.
...
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.

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

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

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

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

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

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

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

RD-0242-HC.
http://www.friends-partners.org/part...s/rd0242hc.htm

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

Bob Clark

**Exoscientist** · 08-05-11, 06:33 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/loo...or-making.html

August 02, 2011

Elon Musk of Spacex talks about a Reusable Falcon Heavy to get to $50 a pound to space.
Two technology areas Musk didn’t like were lifting bodies/wings and nuclear rockets.
On the former, he said he was a “vertical takeoff, vertical landing” type guy and eschewed wings since they had to be tailored for each planet’s atmosphere and were useless on airless bodies such as the Moon.
Drawbacks to nuclear power included the need for shielding (heavy), water (heavy), and public objections against launching nuclear fuel on a rocket. “It’s a tricky thing getting a reactor up there with a ton of uranium,” Musk said and went on to say while nuclear power would be useful for Mars or lunar operations, he implied that some assembly (i.e., mining and processing fuel off planet) would be required.

{emphasis added - B.C.}
http://nextbigfuture.com/2011/08/elo...lks-about.html

Bob Clark

**Exoscientist** · 08-13-11, 12:06 AM

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

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

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

3.)Merlin (rocket engine)
4 Merlin 1C Engine specifications
http://en.wikipedia.org/wiki/Merlin_...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/...thane_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_...#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.s...33c95a22?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/comparat...cket-and/15354

13.)Reusable Launch System.
http://en.wikipedia.org/wiki/Reusabl...zontal_landing

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

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

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

17.)A Study of Air Launch Methods for RLVs.
Marti Sarigul-Klijn, Ph.D. and Nesrin Sarigul-Klijn, Ph.D.
AIAA 2001-4619
p.13
http://mae.ucdavis.edu/faculty/sarig...a2001-4619.pdf

**Exoscientist** · 12-08-11, 08:43 AM

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