An SSTO as "God and Robert Heinlein intended".

[RGClark]

The SpaceLaunchReport.com site operated by Ed Kyle provides the
specifications of some launch vehicles. Here's the page for the Falcon
1:

Space Launch Report: SpaceX Falcon Data Sheet.
http://www.spacelaunchreport.com/falcon.html

Quite interesting is that the total mass and dry mass values for the
Falcon 1 first stage with Merlin 1C engine give a mass ratio of about
20 to 1. This is notable because a 20 to 1 mass ratio is the value
usually given for a kerosene-fueled vehicle to be SSTO. However, this
is for the engine having high vacuum Isp ca. 350 s. The Merlin 1C with
a vacuum Isp of 304 s probably wouldn't work.
However, there are some high performance Russian kerosene engines that
could work. Some possibilities:

Engine Model: RD-120M.
http://www.astronautix.com/engines/rd120.htm#RD-120M

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

Engine Model: RD-0234-HC.
http://www.astronautix.com/engines/rd0234.htm

However, I don't know if this third one was actually built, being a
modification of another engine that burned aerozine.

Some other possibilities can be found on the Astronautix site:

Lox/Kerosene.
http://www.astronautix.com/props/loxosene.htm

And on this list of Russian rocket engines:

Russian/Ukrainian space-rocket and missile liquid-propellant engines.
http://www.b14643.de/Spacerockets_1/Diverse/Russian%2520engines/engines.htm

The problem is the engine has to have good Isp as well as a good T/W
ratio for this SSTO application. There are some engines listed that
even have a vacuum Isp above 360 s. However, these generally are the
small engines used for example as reaction control thrusters in orbit
and usually have poor T/W ratios.
For the required delta-V I'll use the fact that a dense propellant
vehicle may only require a delta-V of 8,900 m/s, compared to a
hydrogen-fueled vehicle which may require in the range of 9,100 to
9,200 m/s. The reason for this is explained here:

Hydrogen delta-V.
http://yarchive.net/space/rocket/fuels/hydrogen_deltav.html

Then when you add on the fact that launching near the equator gives
you 462 m/s for free from the Earth's rotation, we can take the
required delta-V that has to be supplied by the kerosene-fueled
vehicle as 8,500 m/s.
I'll focus on the RD-0124 because of its high Isp, 359 s vacuum and
331 s sea level. On the "Russian/Ukrainian space-rocket and missile
liquid-propellant engines" page its sea level thrust is given as
253,200 N, 25,840 kgf. However, the Falcon 1 first stage weighs 28,553
kg. So we'll need two of them. Each weighs 480 kg, so two would be 960
kg. This is 300 kg more than the single Merlin 1C. So the dry mass of
the Falcon 1 first stage is raised to 1,751 kg. There is a RD-0124M
listed on the Astronautix page that only weighs 360 kg, but its sea
level Isp and thrust are not given, so we'll use the RD-0124 until
further info on the RD-0124M is available.
Taking the midpoint value of the Isp as 345 s we get a delta-V of
345*9.8ln(1 + 27102/1751) = 9,474 m/s (!) Note also the achieved delta-
V would actually be higher than this because the trajectory averaged
Isp is closer to the vacuum value since the rocket spends most of the
time at altitude.
This calculation did not include the nose cone fairing weight of 136
kg. However, the dry mass for the first stage probably includes the
interstage weight, which is not listed, since this remains behind with
the first stage when the second stage fires. Note then that the
interstage would be removed for the SSTO application. From looking at
the images of the Falcon 1, the size of the cylindrical interstage in
comparison to the conical nose cone fairing suggests the interstage
should weigh more. So I'll keep the dry mass as 1,751 kg.
Now considering that we only need 8,500 m/s delta-V we can add 636 kg
of payload. But this is even higher than the payload capacity of the
two stage Falcon 1!
We saw that the thrust value of the RD-0124 is not much smaller than
the gross weight of the Falcon 1 first stage. So we can get a vehicle
capable of being lifted by a single RD-0124 by reducing the propellant
somewhat, say by 25%. This reduces the dry weight now since one
RD-0124 weighs less than a Merlin 1C and the tank mass would also be
reduced 25%. Using an analogous calculation as before, the payload
capacity of this SSTO would be in the range of 500 kg.
We can perform a similar analysis on the Falcon 1e first stage that
uses the upgraded Merlin 1C+ engine. Assuming the T/W ratio of the
Merlin 1C+ is the same as that of the Merlin 1C, the mass of the two
of the RD-124's would now be only 100 kg more than the Merlin 1C+.
The dry mass and total mass numbers on the SpaceLaunchReport page for
the Falcon 1e are estimated. But accepting these values we would be
able to get a payload in the range of 1,800 kg. This is again higher
than the payload capacity of the original two stage Falcon 1e. In fact
it could place into orbit the 1-man Mercury capsule.
The launch cost of the Falcon 1, Falcon 1e is only about $8 million -
$9 million. So we could have the first stage for that amount or
perhaps less since we don't need the engines which make up the bulk of
the cost. How much could we buy the Russian engines for? This article
says the much higher thrust RD-180 cost $10 million:

From Russia, With 1 Million Pounds of Thrust.
Why the workhorse RD-180 may be the future of US rocketry.
Issue 9.12 | Dec 2001
"This engine cost $10 million and produces almost 1 million pounds of
thrust. You can't do that with an American-made engine."
http://www.wired.com/wired/archive/9.12/rd-180.html

This report gives the price of the also much higher thrust AJ26-60,
derived from the Russian NK-43, as $4 milliion:

A Study of Air Launch Methods for RLVs.
Marti Sarigul-Klijn, Ph.D. and Nesrin Sarigul-Klijn, Ph.D.
AIAA 2001-4619
"The main engine is currently proposed as the 3,260
lb. RP-LOX Aerojet AJ26-60, which is the former
Russian NK-43 engine. Thrust to weight of 122 to
1 compares to the Space Shuttle Main Engine’s
(SSME) 67 to 1 and specific impulse (Isp = 348.3
seconds vacuum) is 50 to 60 seconds better than
the Atlas II, Delta II, or Delta III RP-LOX engines.
A total of 831 engines have been tested for
194,000 seconds. These engines are available for
$4 million each, which is about 10% the cost of a
SSME."
http://mae.ucdavis.edu/faculty/sarigul/aiaa2001-4619.pdf

Then the much lower thrust RD-0124 could quite likely be purchased
for less than $4 million. So the single RD-0124 powered SSTO could be
purchased for less than $12 million.

Even though the mathematics says it should be possible, and has been
for decades, it is still commonly believed that SSTO performance with
chemical propulsion is not possible even among experts in the space
industry:

Space Tourism is a Hoax
By Fredrick Engstrom and Heinz Pfeffer
11/16/09 09:02 AM ET
"In 1903, the Russian scientist Konstantin Tsiolkovsky established the
so-called rocket equation, which calculates the initial mass of a
rocket needed to put a certain payload into orbit, given that the
orbital speed is fixed at 28,000 kilometers per hour, and that the
maximum speed of the gas exhausted from the rocket that propels it
forward is also fixed.
"You quickly find that the structure and the tanks needed to contain
the fuel are so heavy that you will never be able to orbit a
significant payload with a single-stage rocket. Thus, it is necessary
to use several rocket stages that are dumped on the way up to get any
net mass, i.e. payload, into orbit.
"Let us look at the most successful rocket on the market — the
European Ariane 5. Its start weight is 750 tons, of which 650 tons are
fuel, 80 tons are structure and around 20 tons are left for low Earth
orbit payload.
"You can have a different number of stages, and you can look for minor
improvements, but you can never get around the fact that you need big
machines that are staged to reach orbital speed. Not much has happened
in propulsion in a fundamental sense since Wernher von Braun’s Saturn
rocket. And there is nothing on the horizon, if you discount
controlling gravity or some exotic technology like that. In any case,
it is not for tomorrow."
http://www.spacenews.com/commentaries/091116-space-tourism-hoax.html

The Cold Equations Of Spaceflight.
by Jeffrey F. Bell
Honolulu HI (SPX) Sep 09, 2005
"Why isn't Mike Griffin pulling out the blueprints for X-30/NASP, DC-X/
Delta Clipper, or X-33/VentureStar? Billions of dollars were spent on
these programs before they were cancelled. Why aren't we using all
that research to design a cheap, reusable, Single-Stage-To-Orbit
vehicle that operates just like an airplane and doesn't fall in the
ocean after one flight?"
"The answer to this question is: All of these vehicles were fantasy
projects. They violated basic laws of physics and engineering. They
were impossible with current technology, or any technology we can
afford to develop on the timescale and budgets available to NASA. They
were doomed attempts to avoid the Cold Equations of Spaceflight."
http://www.spacedaily.com/news/oped-05zy.html

Then it is important that such a SSTO vehicle be produced even if
first expendable to remove the psychological barrier that it can not
be done. Once it is seen that it can be done, and in fact how easily
and cheaply it can be done, then there it will be seen that in fact
the production of SSTO vehicles are really no more difficult than
those of multistage vehicles.
Then will be opened the floodgates to reusable SSTO vehicles, and low
cost passenger space access as commonplace as trans-oceanic air
travel.


Bob Clark



---------- Post added at 02:20 PM ---------- Previous post was at 02:18 PM ----------

...
Even though the mathematics says it should be possible, and has been
for decades, it is still commonly believed that SSTO performance with
chemical propulsion is not possible even among experts in the space
industry...
Then it is important that such a SSTO vehicle be produced even if
first expendable to remove the psychological barrier that it can not
be done. Once it is seen that it can be done, and in fact how easily
and cheaply it can be done, then there it will be seen that in fact
the production of SSTO vehicles are really no more difficult than
those of multistage vehicles.
Then will be opened the floodgates to reusable SSTO vehicles, and low cost passenger space access as commonplace as trans-oceanic air travel.

The calculation of the Falcon 1 first stage with more efficient engines having SSTO capability leads me to a surprising conclusion: it won't even have to be millionaires who could own such SSTO's.
For instance to "own", in the sense of live in, a million dollar home you don't have to have a million dollar income or even a million dollar net worth.
You just have to make the mortgage payments, which per year can be a fraction of the million dollar cost of the home. This is in the salary range of many just upper class individuals. Then such orbital rockets with financing will be in the cost range of many such individuals.
A combination of factors suggest this is possible. First, with mass production the cost of the rocket structure and of the engines will drop significantly. Also, though the Falcon 1 is priced at about $8 million, remember a large proportion of this is to cover development cost. The large majority of this cost was for the development of the engines. But neither of these two SpaceX engines would be used for the SSTO purpose. Instead would be used the much cheaper for their size Russian engines.
We can estimate how much they would cost based on their size and the costs for much larger, i.e., more powerful, Russian engines. The 1,000,000 lbs. thrust RD-180 costs $10 million. The 400,000 lbs. thrust NK-33 costs $4 million. Based on this we can estimate the cost of the 60,000 lbs. thrust RD-0124 I was recommending for the SSTO purpose as $600,000.
Another reason for why such SSTO's will be lower cost than the $8 million Falcon 1 is that the manufacturing cost is actually only a fraction of the launch cost. See for example the estimates in Tables 3 and 4 here:

When Physics, Economics, and Reality Collide. The Challenge of Cheap Orbital Access.
by John M. Jurist, M.D., Sam Dinkin, Ph.D, David Livingston, DBA
http://www.colonyfund.com/Reading/papers/phys_econ_leo.html#elv

Note then the methods for achieving such high mass ratios as with the Falcon 1 first stage don't appear to be especially hard. See for example the description of the Falcon 1 propellant tanks given here:

Falcon 1 Overview.
http://www.spacex.com/falcon1.php#first_stage

They appear to use a combination of methods known for decades such as a common bulkhead and an isogrid design. So using these methods, similar high mass ratios could easily be achieved by other aerospace companies. Actually little in research and development costs would be required for the structures.
There is another key cost that figures into launch costs mentioned in the "When Physics, Economics, and Reality Collide: The Challenge of Cheap Orbital Access" article. That is the cost of range access, usually provided by governments. With wide numbers of privately owned rockets launching daily this cost could be reduced significantly.
A big component of the research and development costs however would be the actual flight tests. This would be significantly reduced with reusable systems.
For this to have a high demand you would need for it to be manned-flight capable. The Falcon 1e sized SSTO would require two RD-0124 engines for a 1,800 kg payload capacity, but would be able to loft a Project Mercury-sized capsule:

Mercury.
http://www.braeunig.us/space/specs/mercury.htm

You would also need a lightweight reentry system. Inflatable heat shields may fit the bill:

NASA Launches New Technology: An Inflatable Heat Shield.
08.17.09
http://www.nasa.gov/topics/aeronautics/features/irve.html

Pod People.
They're the ones thinking outside the space capsule.
* By James Oberg
* Air & Space Magazine, November 01, 2003
"IN 1964, MOST VIEWERS OF TELEVISED SPACE "SHOTS," AS THEY WERE CALLED THEN, knew what it took to protect a spacecraft from the fire of reentry. It took big, heavy shields bolted to pressurized metal vessels. One of the most nerve-racking moments of the early space program had been the final minutes of John Glenn’s 1962 Mercury flight, when Mission Control waited to learn whether his shield had remained attached to the Friendship 7 capsule during the violent return.
"Two years later, on June 10, 1964, another, much lighter vehicle entered the atmosphere with no one on board. In engineering terms it was nearly as daring as the Mercury flights had been. Launched on a sounding rocket to an altitude of 96 miles over New Mexico, the craft dove back toward Earth at a speed of more than 5,000 mph. Being so light, it didn’t generate as much heat from atmospheric friction as Glenn’s capsule had, so it had only a thin coating of thermal protection—no shield. Odder still, it was inflated like a balloon in a Thanksgiving day parade."
http://www.airspacemag.com/space-exploration/cit-oberg.html

This second, which involves a lifting body inflatable heat shield, would result in significant reduction in reentry heating by making use of shapes optimized for high lift/drag ratio at hypersonic speed.


Bob Clark

---------- Post added at 02:22 PM ---------- Previous post was at 02:20 PM ----------

I'm looking for a numerical trajectory integration program.
These are the conditions under which I want to estimate the required delta-V to orbit:

1.)use a dense propellant such as kerosene/LOX; dense propellants are known to reduce gravity losses.

2.)use a moderate to high liftoff thrust/weight ratio, say, 1.4 and above; high liftoff T/W also reduces gravity losses.

3.)launch near equator to get the ca. 460 m/s tangential boost.

4.)only get to 100 km, the altitude considered space, to just launch satellites or make orbital transfers, not for long term orbits.

Anyone know if the Orbiter sim could do this?

Bob Clark

 

[Sky Captain]

I suggest you just take a suitable Velcro rockets stage since it is config file based and easy to edit. Set the physical values from Falcon 1 first stage and test fly it in Orbiter. I think it should work with reasonable accuracy.

While an expendable SSTO booster might be relatively easy to achieve a reusable one would be a completely different matter. Suppose a Falcon 1 based SSTO has 600 kg payload. To make it reusable would need a thermal protection, aerodynamic surfaces and some sort of recovery system to for gentle landing which would quickly add extra mass.

 

[Urwumpe]

Gravity losses depend on trajectory and burn time, not propellant density. This is a fallacy. Only because most rockets with high propellant density also have a short burn time (solids!), this can backfire for SSTOs. Assuming 8900 m/s as first-order estimate for designing an SSTO is exactly the wrong direction. First of all, because the value is way off reality, it is calculated by assuming an SLBM ascent profile (no acceleration restraints, extremely high thrust), which needs only little over two minutes for reaching MECO - yes, there you have really low gravity losses.

If you also want to reduce acceleration loads on the payload and limit to 3G, you will quickly approach 300 seconds burn time as practical minimum. There is no engineering law cast in silicon carbide, that says that cryogen rocket engines can't provide 3G acceleration for most of the flight.

Also, you need to include aerodynamic loads on the rocket - you CAN use brute force for getting through a high max-Q, but this increases your structural mass, resulting in mass that you will have to carry with you until landing. So, unless you can afford the higher structural mass, you will want to have net acceleration down to 1.2 g or even less during the Max-Q phase, and less than 2g on lift-off, because you otherwise have max-Q at too low altitude and too high air density.

There is a reason why we still have no SSTO - TANSTAAFL. On the paper, it is simple, if you can choose your first-order estimates to make things work out. But once you get into further iterations of the design, you will be forced to correct your optimistic first-order estimates, making your SSTO quickly unrealistic. And that way is standard for all SSTOs, likely because the first order estimate makes you select a design, that is optimal for your optimistic estimates, but not surviving in reality.

 

[Sky Captain]

What if you drop engines to cut down extra mass and also to limit acceleration? IIRC there was a version of Alas rocket in the sixties which only droppped two engines but otherwise reached orbit like SSTO on a single stage. When considering expandable SSTO's there is no reason to bring all engines to orbit.

 

[Ghostrider]

Dropping engines like the Atlas means more complex tubing because you need to cut off and isolate the fuel lines that go to the jettisoned engines. Atlas was made that way because it was still unknown if a rocket engine could be ignited in space, so the safest thing was to light them up all of them at lift-off. When it was proven that you could indeed ignite an engine in space, the concept was abandoned. That staging procedure was also the cause of many a problem with the Ranger moon probes.

 

[Urwumpe]

Additionally: any staging system also means a mass penalty. It is not much mass, but small masses sum up to big mass increases in your design - see the Ares family as example, that not only American women quickly gain weight.

And staging looks bad on a SSTO.

The engines of the Atlas also had not been little mass - if I remember correctly, they reduced the structural mass by 40%.

 

[Andy44]

Atlas (the real Atlas) was an engineering marvel. Unless I'm mistaken it remains the rocket with the best mass-ratio. Starting all the engines on the ground was indeed designed in from the beginning due to development risk for air-starting, but by the time the program was in production air-starting was pretty far along. Thus the Agena and Centaur upper stages later used with the Atlas.

And as Urwumpe stated, the two booster engines were heavy beasts and jettisoning them made the system work. The fuel tank was very lightweight and used pressurization to achieve strength with low mass. The design was among the earliest rockets but it was so successful it flew unti 2004.

Convair even studied the idea of a fly-back Atlas during all the space shuttle studies.

 

[Urwumpe]

Also, the balloon tanks of Atlas also had been its major weakness - it was impossible to handle the stage on the ground without pressurizing the tank and use special stretchers for the stage to keep it under tension of the pressure is too low. Otherwise, the rocket would just collapse under its own weight.

 

[Sky Captain]

The Space shuttle external tank weigh 26,5 tons empty and carry 735 tons of fuel which gives it a mass ratio of 27 to 1. It would require six SSME's to reliably get off the ground. Six SSME's is 19 tons of mass which would make mass ratio 16 to 1 still plenty to achieve orbit. Obviously ET would require redesign to accept thrust from the bottom, but structurally thrust from the bottom is better than from the side like it is now so it might be possible to make ET a bit more lighter. Overall it should be possible to turn ET into SSTO booster however throwing away six SSME's on every launch would make it the most expensive rocket in the world.

 

[Urwumpe]

Add a massive thrust structure at the bottom, to transport the thrust forces to the tank structure (engines don't work well floating in the air). Add hydraulic system for gimballing the engines. Add guidance system, which is a small mass, but needed.

What you want to do is basically a low-performance version of the Advanced Launch System, which NASA studied in the early 1990s.

http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19910018894_1991018894.pdf

 

[Sky Captain]

But ET dry mass already include structural strengthening to accept thrust from SRB's and shuttle. In SSTO application those are not needed so you would basically remove now unnecessary thrust structures from sides and build a new thrust structure at the bottom. Since in this application there are no SRB's that apply huge loads to the ET the upper part could probably made lighter than it is now. Even if after redesign ET with six SSME's have mass ratio of 10 to 1 it would still reach orbit and could carry some payload. I' m not trying to argue that it would be cost effective way to launch something just to prove a point that it should be possible to build an expandable SSTO with existing technology.

 

[Urwumpe]

Yes - some payload. Will be pretty tiny, and still you have only managed to reach orbit without returning to Earth again, which makes SSTO a pure bonus. If you can ignore the ambient pressure effects of the specific impulse, you will get just 3.24% payload mass in your 0.1 structural mass constant plan. With ambient pressure, it would even be negative.

Also you need to add RCS/OMS anyway, if you want to be able to have a accurate orbit insertion. Again, additional mass.

And don't overestimate the savings in mass by leaving side-mount and SRBs away: The SRBs also provide a large amount of stability to the ET during the early maneuvers. Most heavy structure is only in the intertank structure of the ET, where the loads of the SRBs are accepted and the thrust inbalances of both SRBs negotiated.

 

[Andy44]

If you were using the ET as a conventional rocket body like that, wouldn't you swap the locations of the LH2 and LOX tanks?

 

[Urwumpe]

If you were using the ET as a conventional rocket body like that, wouldn't you swap the locations of the LH2 and LOX tanks?

Yes, you would prefer to do so, so the denser fuel is aft and thus the CoG closer to the engines. But that is not too critical, since rockets are always instable, you just choose HOW instable it is.

 

[Andy44]

According to the History of STS: First 100 Missions book, the ET had to gain weight to put the heavier LO2 in the nose for stability purposes, and that if they could've put the LO2 in the tail then it would've made for a lighter ET.

Sounds to me like building a SSTO version of ET/SSME really means building a whole new ET. And while you're at it, build a high-chamber pressure expendable engine that's cheaper than the SSME, too. SSMEs only really make sense if you're recovering them.

 

[Urwumpe]

AFAIR, the top-heaviness of the ET is also required because of the side-mount, otherwise the location of the SSMEs would be far worse.

 

[Sky Captain]

In SSTO application you could also eliminate the intertank structure and choose a design with common bulkhead between LOX and LH2 tanks to further reduce weight. Also you could eliminate the pointy nose and make the front section standard cylindrical shape to save even more weight and make putting a payload on it easier.

 

[MikeB]

Disclaimer: I haven't studied the details. Does anyone here think that a Skylon-like SSTO spacecraft is realistically feasible? The use of air-breathing engines in the air seems like a good idea, even if the specific design of the Sabre engine doesn't work out.

 

[Urwumpe]

Disclaimer: I haven't studied the details. Does anyone here think that a Skylon-like SSTO spacecraft is realistically feasible? The use of air-breathing engines in the air seems like a good idea, even if the specific design of the Sabre engine doesn't work out.

The Skylon is pretty feasible, even the Sabre engines are not unrealistic, just in the beginning of their development. The design of the Sabre works in the math, the worse problem will be getting the thermal insulation of the hydrogen tanks light enough. They have a spreadsheet on their homepage, with the ascent/descent trajectories and the calculated numbers, they look pretty sound to me.

 

[RGClark]

I suggest you just take a suitable Velcro rockets stage since it is config file based and easy to edit. Set the physical values from Falcon 1 first stage and test fly it in Orbiter. I think it should work with reasonable accuracy.

While an expendable SSTO booster might be relatively easy to achieve a reusable one would be a completely different matter. Suppose a Falcon 1 based SSTO has 600 kg payload. To make it reusable would need a thermal protection, aerodynamic surfaces and some sort of recovery system to for gentle landing which would quickly add extra mass.

Thanks for the info. I haven't used the Orbiter sim yet. How long would it take to get up to speed to do this?

Bob Clark