An SSTO as "God and Robert Heinlein intended".

[RGClark]

To be honest that sort of conversion doesn't make any sense to me; I would rather build an entirely new vehicle, it can be based (for example aerodynamics wise) on something like the X-37B, but it trying to shoehorn a totally different goal into the same airframe would probably be very inefficient.

Since it is twice the size of the X-37B, it would certainly not be called the X-37B. But we will have a very good idea what the aerodynamic properties will be like from the X-37B example by using similar design but just scaled up.
Remember the purpose of the airframe on the X-37B is just to glide back, which is the same purpose it will be used for with the SSTO version.


Bob Clark

 

[RGClark]

still not enough. You need something with about triple the energy density of Hydrazine. That is not just a higher specific impulse, but also a higher density of the fuel as hydrazine, and that is not kerolox. At the same density, you would need 5400 Ns/kg specific impulse, which is currently not possible. For getting into the range of kerolox, with 4300 Ns/kg at vacuum (using a high-pressure, high expansion engine), you would need 25% higher density of kerolox over hydrazine. and hydrazine has roughly the density of water.

Perhaps you are thinking of a vehicle of the same size as the X-37B? It is known that if you can scale up a rocket you can improve your mass ratio.
To have a SSTO, it is commonly said that for a hydrogen-fueled version you need a mass ratio of 10 to 1 and for a kerosene version you need a mass ratio of 20 to 1.
It is important to remember also I am filling pretty much the entire internal volume with propellant, no payload bay or even avionics bay. So it's not just a matter of density of the propellant since you actually have more volume to hold the propellant.


Bob Clark

 

[Urwumpe]

Forget this idea quickly. A spacecraft is no aircraft, in which you can fill every volume with unpressurized fuel. Also payload or avionics bay are not optional. The payload bay give you the purpose, the avionics the ability. You also can't make them infinitely small or put them anywhere. The avionics for example must be installed in a way, that inertial navigation platform and star trackers or GPS can work together.

 

[RGClark]

Forget this idea quickly. A spacecraft is no aircraft, in which you can fill every volume with unpressurized fuel. Also payload or avionics bay are not optional. The payload bay give you the purpose, the avionics the ability. You also can't make them infinitely small or put them anywhere. The avionics for example must be installed in a way, that inertial navigation platform and star trackers or GPS can work together.

You can certainly fill the payload bay with propellant. The X-33 and the VentureStar in their latest versions were expected to have payload cannisters on top of the vehicle rather than internally.
With modern miniaturization that space that needs to be taken up by avionics is rather small. For the X-37B you need other equipment besides the avionics in the "equipment bay" because of all the different missions the X-37B was expected to do in space. For one thing you need a solar panel for the long space mission.


Bob Clark

 

[Urwumpe]

of course, the external payload canisters had been last-resort solutions to keep the project at least slightly viable. Practically, such canisters are even on simpler shaped rockets a vicious thing.

Also such canisters can only be used for satellites, for transporting experiments, they are less useful, since you can't return them to earth. Extremely bad for those missions, in which you need the wings as well.

 

[T.Neo]

For one thing you need a solar panel for the long space mission.

You need a power source for any space mission. Considering than an SSTO is going to spend at least a day in space if delivering satellites, and multiple days in space if delivering people to a space station or waiting orbital craft (even if it uses only external power while docked), you will need to have a way to generate power... none of the options are zero-mass or zero-volume...

 

[RGClark]

In doing some background web searches, I found that the upper stage of the Direct team's Jupiter-246 vehicle also would become SSTO when switched to a SSME engine. I guess I should not have been surprised by this. The thesis I have been arguing repeatedly via email with individuals in NASA and the industry and on space forums such as this one is that if you use BOTH the most weight optimized designs AND the highest efficiency engines available, then what you will wind up with will be SSTO capable whether you intend it to or not.
...
So exchanging these for a SSME will add about 1,300 kg to the upper stage weight. The dry mass will be increased then to 13,150 kg, and the gross mass to 204,000 kg. However, by the Space Shuttle main engine thrust specifications, even at 109% thrust this comes to only 417,300 lbs, or 189,700 kgf. So we'll reduce the propellant load to be lifted by the SSME...

You could get the SSTO without reducing propellant by using two SSME engines. For a manned launcher this would be preferred to have engine out capability. The dry mass with one SSME I calculated to be 13,150 kg. Adding on a second SSME would bring the dry mass to about 16,300 kg.
The propellant load is 190,850 kg. Then you could loft 7,200 kg payload:

425*9.8ln(1 + 190,850/(16,300 + 7,200)) = 9,210 kg.

Again with propellant densification and recent lightweighting techniques this payload might be raised to about 10,000 kg.



Bob Clark

 

[RGClark]

In doing some background web searches, I found that the upper stage of the Direct team's Jupiter-246 vehicle also would become SSTO when switched to a SSME engine. I guess I should not have been surprised by this. The thesis I have been arguing repeatedly via email with individuals in NASA and the industry and on space forums such as this one is that if you use BOTH the most weight optimized designs AND the highest efficiency engines available, then what you will wind up with will be SSTO capable whether you intend it to or not.
...
We'll take the liftoff thrust/weight ratio to be 1.2. This will bring the gross mass down to 170,000 kg. Then the propellant mass has to be reduced by 34,000 kg. This brings the propellant mass down to 156,850 kg. Note this results in a mass ratio close to 13, well sufficient for SSTO with a hydrogen-fueled engine.
...
Then using Hudson's 425s average Isp for the SSME and the 9,200 m/s required delta-V value for orbit, this stage as an SSTO could loft 6,200 kg to orbit:

425*9.8ln(1 + 156,850/(13,150 + 6,200)) = 9,200 m/s.

Again, we might be able to loft 10% greater total mass to orbit with propellant densification by subcooling and also shave 10% off the structural mass of the stage with the recent weight saving research. This will bring the payload mass up to about 9,000 kg...

There might still be some resistance to using the upper stage as a SSTO. However, the point of the matter is even if you use these upper stages as part of a multistage system you are still better off using both highly weight optimized structures and engines of highest surface-to-orbit-efficiency (not just vacuum optimized engines) at the same time.
We'll use in this case parallel staging of the same sized stages, a bimese version, but using cross-feed fueling. This is a fueling method that has both stages firing, as with parallel staging, but all the propellant is coming from only a single stage at a time. Then when that stage exhausts its propellant, it is jettisoned, and the remaining stage proceeds on with its own full tank of propellant still remaining.
Let's see how much payload we can carry in this case. Again assume the 425s trajectory averaged Isp of Hudson, and the 9,200 m/s required delta-V for orbit.
Estimate the possible payload as 29,000 kg. For the first segment of the flight the achieved delta-V would be: 425*9.8ln(1+156,850/(2*13,150 + 156,850 +29,000)) = 2,305 m/s.
For the second segment, use the 455s vacuum Isp of the SSME's:
455*9.8ln(1 + 156,850/(13,150 + 29,000)) =6,921. And the total delta-V is 9,226 m/s, sufficient for orbit with a 29,000 kg payload.
Note that a 29,000 kg payload is sufficient to even carry a Orion capsule, at least in an expendable version of the staged vehicle without reentry and landing systems.
Then you have different options for the vehicle. As a single stage it could carry a small capsule such as the SpaceX Dragon, or the Boeing CST-100. But using twinned copies of it, it would be able to loft the heavier Orion spacecraft.


Bob Clark

 

[RGClark]

In doing some background web searches, I found that the upper stage of the Direct team's Jupiter-246 vehicle also would become SSTO when switched to a SSME engine. I guess I should not have been surprised by this. The thesis I have been arguing repeatedly via email with individuals in NASA and the industry and on space forums such as this one is that if you use BOTH the most weight optimized designs AND the highest efficiency engines available, then what you will wind up with will be SSTO capable whether you intend it to or not.
By highest efficiency engines I don't mean just an engine optimized to have a high vacuum Isp only. I mean an engine of highest efficiency over the entire flight range to orbit. For hydrogen engines that is the SSME, and the Russian analogue RD-0120...

Another highly weight optimized stage was the S-II second stage on the Saturn V. According to this Wikipedia page it was even better optimized than the S-IVB stage :

Saturn V.
S-II second stage.
http://en.wikipedia.org/wiki/Saturn_V#S-II_second_stage

The 5 J-2 engines used had a mass of 1,580 kg each, for a total mass of 7,900 kg. You'll need 3 of the SSME's operating at 109% thrust to lift the mass. So the 7,900 kg mass of the engines is replaced with 9,300 kg. And the 36,000 kg S-II dry mass is raised to 37,400 kg and the gross mass is raised to 481,400 kg.
Now using Gary Hudson's 425s trajectory averaged Isp for the SSME engines, and the 9,200 m/s required delta-V to orbit. We get a 17,000 kg payload:

425*9.8ln((481400 + 17000)/(37400 + 17000)) = 9,225 m/s

However, again we can get 10% greater total mass to orbit by propellant densification. This brings the payload to 22,440 kg. Also perhaps 10% off the structural mass can be saved by using aluminum-lithium alloy. And an additional 10% mass can be saved by the new weight saving methods. These weight savings can go to extra payload to bring the payload mass up to 28,000 kg. Note this is sufficient now to carry the Orion spacecraft as a SSTO.

Bob Clark

 

[RGClark]

Nice article here arguing in favor of NASA promoting small manned commercial vehicles:

Human spaceflight for less: the case for smaller launch vehicles,
revisited.
by Grant Bonin
Monday, June 6, 2011
http://www.thespacereview.com/article/1861/1

In the comments section, I commented that the capability to produce
such small, low cost, manned vehicles exists now. I estimated the cost
for a such a reusable vehicle in the range of a few tens of millions
of dollars unit cost, comparable to a medium sized business jet.


Bob Clark

 

[RGClark]

Perhaps you are thinking of a vehicle of the same size as the X-37B? It is known that if you can scale up a rocket you can improve your mass ratio.
To have a SSTO, it is commonly said that for a hydrogen-fueled version you need a mass ratio of 10 to 1 and for a kerosene version you need a mass ratio of 20 to 1.
It is important to remember also I am filling pretty much the entire internal volume with propellant, no payload bay or even avionics bay. So it's not just a matter of density of the propellant since you actually have more volume to hold the propellant.


Bob Clark

I just saw this on Hobbyspace.com:


Boeing proposes SSTO system for AF RBS program.
The new issue of Aviation Week has a brief blurb about a Boeing proposal for the Air Force's Reusable Booster System (RBS) program: Boeing Offers AFRL Reusable Booster Proposal - AvWeek - June.13.11 (subscription required).

Darryl Davis, who leads Boeing's Phantom Works, tells AvWeek that they are proposing a 3-4 year technology readiness assessment that would lead up to a demonstration of a X-37B type of system
but would be smaller. Wind tunnel tests have been completed. Davis says the system would be a single stage capable of reaching low Earth orbit and, with a booster, higher orbits. The system would return to Earth as a glider.
Davis says "that advances in lightweight composites warrant another look" at single-stage-to-orbit launchers.
http://www.hobbyspace.com/nucleus/index.php?itemid=30110

I don't have a subscription to AV Week. If anyone does perhaps they could look it up.
I'm curious about the statement it would be "smaller" than the X-37B. I did some preliminary calculations that if you switched to kerosene fuel and a high efficiency engine such as the NK-33, and filled every scrap of internal volume with fuel, then a vehicle twice the size of the X-37B could be SSTO. I'm surprised they are able to get it to work with a smaller vehicle than the X-37B.


Bob Clark

 

[RGClark]

I just saw this on Hobbyspace.com:

Boeing proposes SSTO system for AF RBS program.
The new issue of Aviation Week has a brief blurb about a Boeing proposal for the Air Force's Reusable Booster System (RBS) program: Boeing Offers AFRL Reusable Booster Proposal - AvWeek - June.13.11 (subscription required).
Darryl Davis, who leads Boeing's Phantom Works, tells AvWeek that they are proposing a 3-4 year technology readiness assessment that would lead up to a demonstration of a X-37B type of system
but would be smaller. Wind tunnel tests have been completed. Davis says the system would be a single stage capable of reaching low Earth orbit and, with a booster, higher orbits. The system would return to Earth as a glider.
Davis says "that advances in lightweight composites warrant another look" at single-stage-to-orbit launchers.
http://www.hobbyspace.com/nucleus/index.php?itemid=30110

I don't have a subscription to AV Week. If anyone does perhaps they could look it up.
I'm curious about the statement it would be "smaller" than the X-37B. I did some preliminary calculations that if you switched to kerosene fuel and a high efficiency engine such as the NK-33, and filled every scrap of internal volume with fuel, then a vehicle twice the size of the X-37B could be SSTO. I'm surprised they are able to get it to work with a smaller vehicle than the X-37B.

About it being smaller, perhaps it means smaller than the booster, the
Atlas V, and X-37B system, as the Atlas V weighs upwards of 300,000 kg.
This X-37B derived SSTO would be analogous to the winged version of
the X-33 submitted by Rockwell. As I have been arguing switching to
dense propellants allows you to produce a small fully orbital vehicle, where
as the hydrogen fueled version of comparable size could only be suborbital.

It is important to keep in mind the failure of the Lockheed X-33 does
not show that SSTO's are impossible. If you look at the details of
that program you see what failed was the attempt to form conformal,
i.e., non-cylindrically shaped tanks out of composites. However, the
other advanced features were progressing nicely such as the aerospike
engines and the metallic shingle thermal protection.

Then note that the other suborbital X-33 versions proposed by
Rockwell and McDonnell-Douglas would use circular-cross section tanks
that would be easy to produce. So the expectation is they would have
worked, and thus have provided impetus to proceed to the full size orbital versions.

Again note though if these versions had been switched to dense,
hydrocarbon fueled then the original X-33 versions themselves would
have become actually fully orbital. I showed that a vertical landing
DC-X styled SSTO could be derived from the Delta Thor first stage in post #25 of this thread.

This shows the McDonnell Douglas version of the X-33 also styled on
the DC-X should also be SSTO when switched to hydrocarbon fueled. Note
that in point of fact the Delta Thor derived SSTO is smaller than the
McDonnell Douglas X-33, yet still manages to be fully orbital when in being hydrocarbon fueled.

Here I'll do a calculation to show that a X-37B scaled up by a factor
of two will become SSTO capable when switched to using high efficiency
kerosene engines, such as the NK-33. This will show that the winged
Rockwell version of the X-33 could also become SSTO capable when
switched to hydrocarbon fueled from hydrogen. Actually this scaled up
X-37B will still be smaller than the Rockwell X-33, so that the
Rockwell X-33 would have an even better mass ratio when switched to
hydrocarbon fueled, so be able to carry better payload.

The dimensions of the X-37B I used I estimated from this image:
http://www.collectspace.com/review/atlas_x37b02-lg.jpg
by comparing to the published length and wing span values of the
vehicle. I estimate the width of the main cylindrical body as 5 ft.,
call it 1.5 m, and the length of the main cylindrical body, not
including the nozzle or conical nose cone, as 19.2 ft., call it 5.8
meters. Then the size scaled up twice will give the main cylindrical
body a width of 3 m and a length of 11.6 m.

For a SSTO every scrap of internal space is valuable to hold
propellant so I'll fill the entire main body with propellant tanks.
Any payload, and any avionics or other equipment will be placed either
in the nose cone or in an external canister. So we have a volume of
pi*(11.6)*(1.5)^2 = 82 m^3 for the tanks. Since kero/LOX has an
approx. 1,000 kg/m^3 density, this gives a propellant mass of about
82,000 kg.

Note this is about the mass of the propellant in the Delta Thor first
stage. So we'll build up this SSTO as we did the DC-X styled version
based on this Delta Thor first stage. We need to give it wings. Just
as for vertical landing where about 10% propellant had to be set aside
for powered landing, approx. 10% also of the landing weight has to be
set aside for wings for a gliding landing:

Reusable launch system.
2.3 Horizontal landing
http://en.wikipedia.org/wiki/Reusable_launch_system#Horizontal_landing

So the added weight here is about the same as for the DC-X style
case. Actually we'll be using composites for the wings which will be
of short, stubby shape, so probably half this would suffice.

Now for the thermal protection. We'll be protecting the bottom of the
body and the wings. The length of the full vehicle from end of the
nozzle to tip of nose cone is published as 8.9 m, and the width we
estimated as 1.5 m. This is an area of 13.35 m^2. For the wings,
estimate from the above linked image a width on each side of 1.5 m
from the cylindrical body, and a length of 2.5 m, and the shape on
each side as roughly triangular. Then the total area for the wings on
each side of the body is (1/2)*3*2.5 = 3.75 m^2. And the total area
that has to be covered for the body and wings is 13.35 + 3.75 m^2 =
17.1 m^2.

However, the vehicle is scaled up by a factor of 2 so the area that
needs to be covered by TPS is 4 times as great so to 4*17.1 = 68.4 m^2.
I'll use the AETB ceramic tiles actually used on the X-37B rather than the
the metallic shingles used on the X-33. These are lighter at about 12 kg/m^2.
This gives a thermal protection mass of 820 kg, significantly higher than
the vertical landing DC-X style case I discussed in post #25. Note though
from principles of scaling when the vehicle is made larger the percentage
of vehicle mass taken up the TPS will become smaller as this is growing
by the square of the dimensions while the mass is growing by the cube.

Still we'll have to get some mass savings to get the dry mass
comparable to the DC-X style case. The X-37B is notable for its
composite design. Then for this SSTO you'll also want to use composite
fuel tanks. These won't have the difficulty of the X-33 composite
tanks because they will be of circular cross-section. Composite tank
manufacturers have gotten tanks half as heavy as standard metal tanks.
For aluminum tanks the weight for kero/LOX tanks is about 1/100th the
mass of propellant, so about 820 kg in this case. Then the composite
tanks would weigh about 410 kg, saving 410 kg off the dry mass.

You can get additional mass savings in the wings as well which will
be composite. The usual estimate of 10% of the landed mass is for
metal wings. Using composites you can shave 40% off this mass, so save
200 kg off the dry mass. This brings the mass down to that of the DC-X style case.

Finally, the landing gear weight will again be estimated as 3% of the
landed weight; so this added mass also is the same as the DC-X case:

Landing gear weight.
http://yarchive.net/space/launchers/landing_gear_weight.html

Note too though with composites half this amount would likely
suffice.

Now we see the dry mass is about the same as the DC-X case so again
would be SSTO capable.


Bob Clark

 

[T.Neo]

How can you fill every single internal space with propellant? The best shape for a propellant tank is a spheroid, or a cylinder with spheroid (or ogive) endcaps. As I understand it the complex propellant tank shape of the X-33 is what caused structural problems.

Unless your hull is directly based around the propellant tank (read: it is the propellant tank), then fitting propellant into every nook and cranny will be difficult and likely be too much trouble for what it is worth.

In addition, not all the volume within a propellant tank can actually be filled with propellant. You have internal stringers for rigidity, baffles and suchlike to control fluid dynamics within the tank, various other infrastructure, and ullage space.

If the avionics is in an external pod, how is it recovered? Is the avionics just thrown away after every mission? What about the mounts for this pod? How does it disrupt airflow?

Same goes for the payload. Is it like a fairing on a conventional launcher? Payload has a nonzero volume, and that fairing will have nonzero mass and nonzero drag.

In addition, how are payloads returned from orbit? Arguably this is the sole capability of a spaceplane that is unseen anywhere else with any other spacecraft, despite its un-use in history.

 

[RGClark]

Here is a description (long) of the proposal to turn the X-33 into a SSTO by switching to hydrocarbon fueled:

A kerosene-fueled X-33 as a single stage to orbit vehicle.
http://sciforums.com/showthread.php?t=99514

Quite key to this proposal is getting the conformal tanks to be of comparable weight of usual cylindrically-shaped tanks. This is discussed in section II. I'm most optimistic about the method that partitions the tanks into several compartments. The problem with tanks without circular-cross section is that they naturally want to balloon out. The strong divider panels of the compartments prevent this from happening. The detailed calculations for this method are presented in post #4 in that thread.

Probably you would want to keep the avionics in the nose cone. The payload would be carried in an external payload bay or in jettisonable canister. See for example this description of the VentureStar:

VentureStar.
http://www.spaceandtech.com/spacedata/rlvs/venturestar_sum.shtml

For the case where you would return something from orbit you would use the attached external payload bay.

Note when I first wrote this I was only concerned with reducing the weight of the tanks that were already there. However, these lightweighting methods if successful would also work to be able to make the entire internal volume to be able to hold propellant. You could hold approx. 50% more propellant that way, so launch approx. 50% more mass to orbit, i.e., dry vehicle mass + payload.


Bob Clark

 

[RGClark]

About it being smaller, perhaps it means smaller than the booster, the
Atlas V, and X-37B system, as the Atlas V weighs upwards of 300,000 kg.
This X-37B derived SSTO would be analogous to the winged version of
the X-33 submitted by Rockwell. As I have been arguing switching to
dense propellants allows you to produce a small fully orbital vehicle, where
as the hydrogen fueled version of comparable size could only be suborbital.
It is important to keep in mind the failure of the Lockheed X-33 does
not show that SSTO's are impossible. If you look at the details of
that program you see what failed was the attempt to form conformal,
i.e., non-cylindrically shaped tanks out of composites. However, the
other advanced features were progressing nicely such as the aerospike
engines and the metallic shingle thermal protection.
Then note that the other suborbital X-33 versions proposed by
Rockwell and McDonnell-Douglas would use circular-cross section tanks
that would be easy to produce. So the expectation is they would have
worked, and thus have provided impetus to proceed to the full size orbital versions.
Again note though if these versions had been switched to dense,
hydrocarbon fueled then the original X-33 versions themselves would
have become actually fully orbital...

The original Lockheed version of the X-33 as hydrogen fueled could only be suborbital. By scaling it up twice the VentureStar was supposed to become orbital.
I showed above in post #72 how a scaled up twice X-37B could become fully orbital. But it is key to note this scaled up, orbit-capable X-37B would still be smaller than the Lockheed X-33 which could not reach orbit as hydrogen fueled. This shows you can get a SSTO vehicle more easily and cheaper as hydrocarbon fueled.
In analogy with the X-33, we might want to get a high Mach suborbital X-37B vehicle first to test the technology before scaling up to the fully orbital version. Then we'll see the X-37B itself can accomplish this as hydrocarbon fueled at a much smaller size and lower cost than the X-33.
In post #72 we saw the twice scaled up X-37B had an internal volume to hold 82,000 kg of kero/LOX. So the X-37B itself can hold 1/8th this at 10,500 kg. (Recall we are filling most of the internal volume with propellant, removing avionic, instrument, and payload bays.)
We'll replace now the low efficiency engine AR-2/3 with a high efficiency kerosene engine. The NK-33 probably will be too heavy at 1,200 kg and also overpowered for the purpose. We'll use the RD-0124. At 480 kg, this weighs 380 kg more than the AR-2/3 but it does have a high sea level and vacuum Isp (but see footnote 1.)
Another possibility might be to use two of the RD-0242-HC. This is much lighter at 120 kg, though of lower thrust so two would be required for a total of 240 kg. It's not clear also if this engine was ever built, being a derivative of an engine originally fueled by hypergolic propellants.
We need now the dry mass of the vehicle. The original NASA X-37 was not to exceed 7,500 lb, 3,400 kg in dry weight. We'll use this value of 3,400 kg for the Air Force's X-37B version. For our purposes it may very well weigh less than this since we don't need the equipment bays and the solar cells. But we'll be using a 380 kg heavier engine so let's make the dry weight 3,800 kg.
We'll use altitude compensation. High performance kerolox engines should be able to get 338.3 s average Isp using altitude compensation. Then our delta-V will be:

338.3*9.8ln(1 + 10,500/3,800) = 4,390 m/s, about Mach 14. This is also well above the required velocity for suborbital flight.


Bob Clark

Footnote 1: I've been informed by email that the RD-0124 is able to get such high sea level and vacuum Isp values, because the values given are from the engine being used in two different configurations. The sea level value is when the engine is being used in a first stage, and has a short nozzle. And the vacuum value is when the engine was being used in a upper stage and has a long nozzle. However, by using altitude compensation we should be able to get values close to these.

 

[RGClark]

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 the 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. The sustainer engine used was the LR89-5 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 Spacelauncereport.com site. Kyle takes the average Isp as lying 2/3rds of the way up from the sea level value to the vacuum value. 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.
This is just 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" 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

---------- Post added 06-26-11 at 01:40 PM ---------- Previous post was 06-25-11 at 04:34 PM ----------

Because SSTO's are controversial I should make the disclaimer that citing the references in the prior post should not be construed as the cited authors endorsing the viewpoint I expressed in that post.

Note also in fact that this SSTO has a very good value for a ratio that I believe should be regarded as a better measure, i.e., figure of merit, than the payload ratio 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.


Bob Clark

 

[RGClark]

The original Lockheed version of the X-33 as hydrogen fueled could only be suborbital. By scaling it up twice the VentureStar was supposed to become orbital.
I showed above in post #72 how a scaled up twice X-37B could become fully orbital. But it is key to note this scaled up, orbit-capable X-37B would still be smaller than the Lockheed X-33 which could not reach orbit as hydrogen fueled. This shows you can get a SSTO vehicle more easily and cheaper as hydrocarbon fueled.
In analogy with the X-33, we might want to get a high Mach suborbital X-37B vehicle first to test the technology before scaling up to the fully orbital version. Then we'll see the X-37B itself can accomplish this as hydrocarbon fueled at a much smaller size and lower cost than the X-33.
In post #72 we saw the twice scaled up X-37B had an internal volume to hold 82,000 kg of kero/LOX. So the X-37B itself can hold 1/8th this at 10,500 kg. (Recall we are filling most of the internal volume with propellant, removing avionic, instrument, and payload bays.)
We'll replace now the low efficiency engine AR-2/3 with a high efficiency kerosene engine. The NK-33 probably will be too heavy at 1,200 kg and also overpowered for the purpose. We'll use the RD-0124. At 480 kg, this weighs 380 kg more than the AR-2/3 but it does have a high sea level and vacuum Isp (but see footnote 1.)
Another possibility might be to use two of the RD-0242-HC. This is much lighter at 120 kg, though of lower thrust so two would be required for a total of 240 kg. It's not clear also if this engine was ever built, being a derivative of an engine originally fueled by hypergolic propellants.
We need now the dry mass of the vehicle. The original NASA X-37 was not to exceed 7,500 lb, 3,400 kg in dry weight. We'll use this value of 3,400 kg for the Air Force's X-37B version. For our purposes it may very well weigh less than this since we don't need the equipment bays and the solar cells. But we'll be using a 380 kg heavier engine so let's make the dry weight 3,800 kg.
We'll use altitude compensation. High performance kerolox engines should be able to get 338.3 s average Isp using altitude compensation. Then our delta-V will be:

338.3*9.8ln(1 + 10,500/3,800) = 4,390 m/s, about Mach 14. This is also well above the required velocity for suborbital flight.


Bob Clark

Footnote 1: I've been informed by email that the RD-0124 is able to get such high sea level and vacuum Isp values, because the values given are from the engine being used in two different configurations. The sea level value is when the engine is being used in a first stage, and has a short nozzle. And the vacuum value is when the engine was being used in a upper stage and has a long nozzle. However, by using altitude compensation we should be able to get values close to these.

Being able to be achieve a delta-V this high suggests it could be used as a first stage. Boeing in their announcement about producing a SSTO version made this in response to the Air Force's Reusable Booster System (RBS) program.
However, the Air Force's focus was only on a two stage system with only the first stage booster being reusable and the upper stage expendable. Then it might be useful for Boeing to adapt the current size X-37B for this first stage booster purpose. This would certainly be cheaper and involve less of the proverbial technical risk than the scaled-up SSTO version.
Using the estimated dimensions I gave above in post #72 I estimate the nose cone as 1.2 meters long. For this first stage booster purpose I'll fill even the nose cone with propellant. I'll approximate the nose cone as an ogive shape. This site has a downloadable Excel spreadsheet that allows you to calculate the volume of an ogive:

http://www.vatsaas.org/rtv/tools/computationtools.aspx#volume

Using this I calculate the volume of the nose cone as about 2 m^3. So the volume we could fill with propellant is now 12.5 m^3, and the amount of propellant about 12,500 kg.
For the expendable upper stage I'll use the Falcon 1 first stage but with the Merlin 1C engine swapped out for a more efficient engine. From released info on the Merlin 1C engine I'll estimate its mass as 600 kg. We'll replace it with the RD-0242-HC engine I mentioned before.
As I said I wasn't sure if this version was ever built. However, there have been cases where hydrocarbon versions were derived from hypergolic fueled engines so this should be doable.
What is the performance we can expect from this engine as hydrocarbon fueled? From the high chamber pressure we can conclude this is a high performance engine. Such Russian engines with vacuum optimized nozzles have gotten in the range of 360 s vacuum Isp. This engine though is listed on the Astronautix site as only having an Isp of 312 s. However, this was for its use as a first stage engine, so this is undoubtedly for a version with a short nozzle for sea level use. As a point of comparison the Merlin 1C engine used for first stages only has a vacuum Isp of 304 s. But the vacuum optimized version which only has a longer nozzle has a vacuum Isp of 342 s.
I've also been informed by email that simulations of this engine using the engine performance program Propep gave it an Isp in the 370's with a vacuum optimized nozzle. Therefore using altitude compensation I'll take the vacuum Isp as 360 s.
The total weight though will be higher than the thrust of the RD-0124 engine used on the X-37B first stage. So we'll use parallel staging with both stages firing together and use cross-feed fueling to allow the upper stage to have its full tank remaining after the stages separate.
Now for the calculation of the delta-V. The upper stage of the Falcon 1 now has a 500 kg lighter engine, so the gross mass on this stage is now 22,500 kg and the dry mass 1,000 kg. The average Isp for the first stage we'll take again as 338.3 s. Estimate the possible payload as 1,700 kg. Then the first stage delta-V is:

338.3*9.8ln(1+12.5/(3.8+22.5+1.7)) = 1,224 m/s, the second stage delta-V is:
360*9.8ln(1+21.5/(1+1.7)) = 7,737 m/s, and the total delta-V is 8,961 m/s, sufficient for orbit.

It is notable as well that with a payload this high we can even add reentry/landing systems to the upper stage to get a fully reusable system and still have significant payload. I made an estimate of 28% of the landing mass for reentry/landing systems in this post:

Some proposals for low cost heavy lift launchers.
http://www.orbiter-forum.com/showthread.php?p=249968&postcount=6

though with modern materials we can probably cut this in half. So we could still have 1,400 kg or more for payload even if fully reusable.


Bob Clark

 

[RGClark]

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

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/space_travel:_the_path_to_human_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 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 [ame="http://en.wikipedia.org/wiki/TransHab"]Transhab[/ame] 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 [ame="http://en.wikipedia.org/wiki/NERVA"]nuclear powered rockets have been understood since the 60's[/ame]. 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

 

[T.Neo]

Space travel the path to human immortality?

That must be the most absurd thing I've read in months.

SSTOs will not drop prices down to $100/kg. This has been explained again and again and again and you have not satisfactorily refuted (or attempted to refute) it.

And no, the 2001 space vision would not have come true... for one single, simple reason: There was no need for it to. No reason to go to the Moon, no reason to go to Mars, not even a reason to build a Station V.

The problem with getting propellant from the Moon is that it is a desolate wasteland; there is a deficiency of infrastructure there, you either have to ship from Earth, or attempt to construct on the Moon from what little you can ship from Earth. Physics is one thing, economics is another. Just because it is easier physics-wise to launch propellant from the Moon does not mean it is a logistically viable large-scale operation (unless, perhaps, DeltaGliders have replaced the family car).

Especially if we consider that launch costs would be magically low.

Furthermore Uranium is present only in low concentrations on the surface.

The "trillions of dollars" thing doesn't matter because, as has been discussed, asteroids are difficult to get to, difficult to mine, and they are common heritage of mankind which prevents private exploitation.

The resources within an entire mountain may be worth trillions of dollars too, but it isn't like someone is going to come along and try to process that mountain in a useful amount of time.

And "having uranium" is only a tiny, tiny part of actually building and operating a nuclear rocket engine, or nuclear propelled vehicle. For example you need to design and build the engine itself, which would be pretty difficult to do from lunar materials. And the exhaust velocity increase that you have with a nuclear rocket is not everything; nuclear propulsion comes with many disadvantages as well.

2001 being the year of the turn of the millennium (no, it was NOT in the year 2000.)

Yeah, it was- depending on which definition you take. Back then they regarded it to be in 2001, by the time 2000 came around everyone was just pretty thrilled that the first digit changed from a '1' to a '2'...

 

[RGClark]

Space travel the path to human immortality?

That must be the most absurd thing I've read in months.

Again, you just read a title without reading what the article was about.

Bob Clark

---------- Post added at 06:31 PM ---------- Previous post was at 06:27 PM ----------

Yeah, it was- depending on which definition you take. Back then they regarded it to be in 2001, by the time 2000 came around everyone was just pretty thrilled that the first digit changed from a '1' to a '2'...

There was no definition that regarded the new millennium as starting in 2000. That came from a misunderstanding of when the new century started and therefore when the new millennium started.
I could say the new year starts on my birthday. That doesn't mean it's valid.


Bob Clark