Nuclear Space Ship SSTO Proposal

This is an excerpt of a very lengthy explanation of what a nuclear SSTO (Single Stage To Orbit) fully reusable rocket would look like. The full article can be found at the link above.

In this section I describe a huge nuclear powered rocket launcher. I will repeat and expand upon many of the points I made above, because I don't want to throw cryptic acronyms around. I want people to understand just how powerful we can make this rocket if we decide to do it.

The most important difference between our new booster and the Saturn V is in the engines. The Saturn V used five massively powerful F1 engines in the first stage, burning kerosene and liquid oxygen. The mighty F1 produced 1.5 million pounds of thrust. Despite its large size and power, the F1 was a very "relaxed" design. It ran well inside the possible performance envelope. The reason it did so was to increase reliability. This is a sound design principle, so I will apply it to the new launcher wherever possible.

For an engine, I will designate a Gaseous Core Nuclear Reactor design, of the Nuclear Lightbulb subvariant. I like the gas core design for a number of reasons, and the nuclear lightbulb variant for several more.

To recap, the efficiency and power of the thruster is based on the difference in temperature between the fissioning mass and the reaction mass. If you run a solid core NTR much above 3000 C, it melts. This provides a firm "ceiling" on how efficient a solid core reactor can be. A gas core design STARTS melted. In addition, since all of the structure of the fuel mass is dynamic, a gas cored reactor is inherently safer than a solid core device. If a "hot spot" develops in a solid core, disaster ensues. If a hot spot develops in a gas core, the hot spot superheats and "puffs" itself out of existence. A gas core reactor is expected to operate at temperatures of 25,000C. The much higher temperature gradient makes the thruster inherently more efficient.

Second, a solid core reactor has a "fixed" core, since it is solid. A gas core reactor does not, and the radioactive fuel is easily "sucked" out of the core and stored in a highly non-critical state completely out of the engine! The fuel storage system I propose is a mass of thick walled boron-aluminum alloy tubing. As I said above, the fuel proper is uranium hexaflouride gas. UF6 is mean stuff, but we have decades of experience handling it in gaseous diffusion plants, and common aluminum and standard seals are available which resist attack from it. It is stoichiometric, fluorine is low activation, and UF6 changes phase at moderate temperatures, allowing it to be converted from high pressure gas to a solid and back again using nothing fancier than gas cooling and electrical heaters. This naturally makes dealing with the engine easier.

In addition, the design of the gas core allows the addition and removal of fuel "on the fly." The core can also have its density varied by control of the vortex, which directly affects criticality. Both of these elements allow very potent control inputs to be applied to a gas core reactor which are very stable and unaffected by the isotopic condition of the fuel mass.

Also, to repeat, due to the extremely high temperature gradient in the motor, the main cooling of the fissioning mass is not conductive but radiative, a mode which is inherently less susceptible to perturbations. (Having no working fluid for cooling means no material characteristics for the working fluid must be considered.) This radiative cooling mechanism is what allows the "lightbulb" system to work. The silica bulb just has to be transparent enough to let the gigantic power output of the fissioning core flow through, while keeping the radioactive material of the core safely contained inside the thruster. No radioactive materials leak out of the exhaust, it is completely "clean."

Third, a gas cored reactor has several potential "scram" modes, both fast and slow, and the speed of the reaction is easily "throttled" by adding and removing fuel or by manipulating the vortex. A 'scram' is an emergency shutdown, usually done in a very fast way. For example: a gas cored reactor can be fast scrammed by using a pressurized "shotgun" behind a weak window. If the core exceeds the design parameters of the window, which are to be slightly weaker than the silica "lightbulb," then the "shotgun" blasts 150 or so kilos of boron/cadmium pellets into the uranium gas, quenching the reaction immediately. A slightly slower scram which is implemented totally differently is to vary the gas jets in the core to instill a massive disturbance into the fuel vortex. This disturbance would drastically reduce criticality in the fission gas. A third scram mode, slightly slower still, is to implement a high-speed vacuum removal of the fuel mass into the storage system. Having three separate scram modes, one of which is passively triggered, should instill plenty of safety margin in the nuclear core of each thruster. Extensive work was done on gas core reactors, and 25 years ago several experimental designs were built and run successfully. There were technical challenges, but nothing that seems insurmountable or even especially difficult given our current computer and material skills.

A Gas cored NTR using a silica lightbulb. The silica bulb is cooled and pressure-balanced against the thrust chamber by high pressure hydrogen gas. The cooling gas from the silica bulb is used to power three turbopumps "borrowed" from the Space Shuttle Main Engine. These pumps are run at a very relaxed 88 percent of rated power at their maximum setting. The three pumps move 178 kilos of liquid hydrogen per second combined. Most of this is sprayed into the thrust chamber. A portion of the liquid hydrogen is forced into cooling channels for the thrust chamber and expansion nozzle, where a portion of it is bled from micropores to form a cooling gas layer. The gaseous hydrogen that is not bled then flows down the silica lightbulb to cool it, and the cycle finally goes into powering the turbopumps.

This engine produces 1,200,000 pounds of thrust, with an exhaust velocity of 30,000 meters per second, from a thermal output of approximately 80 gigawatts. This equates to an Isp of 3060 seconds. Several sources state that a gas core NTR can exceed 5000 seconds Isp, so 3060 is well inside the overall performance envelope. The three turbopumps from the SSME are run at low power levels, and even losing a pump allows the engine to continue running as long as there is no damage to the nuclear core. Lets assume this design is able to achieve a thrust to weight ratio of ten to one, so the engine and all of its safety systems, off-line fuel storage, etc, weighs 120,000 pounds. I think we can build this engine easily for 60 tons.

Since we are using the Saturn V as our template, we will make the new machine about the same weight, or six million pounds launch weight. With our engines giving 1.2 million pounds of thrust, we need at least five to get off the ground. But, since we have the power of nuclear on our side, we will use seven engines instead of five. Why seven? The most vulnerable moments of a rocket launch are the first fifteen seconds after launch. If we have to scram a motor in those fifteen seconds, having two extras is very comforting. Engine failures further along the flight profile are much easier to recover from, and having two spare engines allows us to be very "chicken" on our criteria for scramming a motor. We can shut one down even at one second after launch if we need to with no risk of crashing the entire vehicle. This further lowers the risk of nuclear power as a means of getting off the earth. With seven engines, we have a thrust of 8.4 million pounds available. In addition, the turbopumps can "overthrottle" the engines easily in dire straits. This gets more thrust at the expense of less Isp.

Let's design the vehicle for a total DeltaV of 15 km per second. This is very high for a LEO booster, but the reason for it is to allow enough reaction mass to perform a powered descent. In other words, this is a true spaceship, that flies up and then can fly back down again.

'M0' is the initial mass of the vehicle, and we have set this to be 6 million pounds.

The value of M1 is what we need to find, since we know we want a total DeltaV of 15,000 m/s.

Doing a little simple math, we find we need 2,400,000 pounds of reaction mass. Since we are using liquid hydrogen, we can now calculate the size of the hydrogen tank needed, which is 15,200 cubic meters. This works out to be a whopping 20 meters in diameter and 55 meters long!

We look at the Saturn V and find our new booster is going to be quite plump compared to the sleek Saturn V, but we have no choice if we want to use liquid hydrogen as reaction mass. Since hydrogen is the best reaction mass physics allows, and is cheap, plentiful, and we have decades of experience handling it, we will use it.

A design height of 105 meters seems reasonable. We assign 15 meters to the engines, 55 meters for the hydrogen tank, 5 meters for shielding and crew space, and a modular cargo area which is 30 meters high and 20 meters in diameter. This is enough cargo space for a good sized office building!

How heavy is the rest of the vehicle? Well, we already decided that the engines are going to weigh 120,000 pounds each, for a total of 840,000 pounds. (To make a comparison, the entire Saturn V, all three stages, engines and all, weighed a mere 414,000 pounds dry.)

Let's splurge here. With nuclear power, we have the power to splurge. Let's use 760,000 pounds to build all of the structure of the new booster. We use thicker and stronger metal, we use extra layers of redundancy, we make it strong and safe and reliable.

We have now used 2,400,000 pounds for reaction mass, 840,000 pounds for the engines, and 760,000 pounds for the rest of the ship's dry structure. This adds up to 4,000,000 pounds, fully built, fully fueled, ready to launch.

But we said at the beginning, the booster has a design launch weight of 6,000,000 pounds! If it only weighs 4 million pounds ready to launch, the rest must be cargo capacity.

Its exhaust is completely clean: It is very difficult to make hydrogen radioactive in a fission reactor. It basically can't happen.

It flies to space with a thousand tons of cargo, and flies back using some gentle aero-braking and its thrusters with another thousand tons of cargo.

This means it has eight times the cargo capacity of the Saturn V, which was not reusable at all. No longer will the Saturn V be the mightiest American rocket. No more resting on our laurels.

With this sort of performance potential, can anyone argue that NTR's are NOT the only sensible course for heavy lift boosters?

There are risks, of course, but careful design and the proper launch site can easily mitigate those risks so that the huge advantages of nuclear propulsion can be realized.

Firstly, horizontal take-off is not more efficient than vertical take-off. Wings only translate the energy you use in thrust into lift, and they translate it badly. Getting low altitude lateral velocity is a waste because all of that is lost in aerodynamic losses at low altitude.

Secondly the 'Cold Equations of Space Flight' article is absolute BS by a fool who doesn't know anything but NASA propaganda. For example, an LH2/LOX SSTO required mass fraction is actually .87, but only if your engines are altitude compensating (like SSMEs and aerospike engines are).

Thirdly, the biggest error in chemical SSTO designs is using LH2 as the sole fuel. LH2 has a terribly low density, which mandates very large fuel tanks and very large aerodynamic cross sections (thus adding lots of aerodynamic losses). Dunn has done a lot of good writing showing that denser fuels like UDMH, methylacetylene, and chilled propane are significantly superior fuels for an SSTO than LH2, to the point of putting as much as two and a half times more payload in orbit.

While the writer is right that air breathing to reduce oxygen load is a right direction to go in, doing so mandates spending lots of time in the 100k-200k altitude range, collecting air, incurring lots of aerodynamic losses. If you are using a low density fuel like LH2, you are wasting your time because the fuel has given your vehicle terrible aerodynamics.

There is an alternative. The RBCC (rocket-based combined cycle) engine that NASA has developed and all but cancelled (what is new, eh?) is the solution. This uses a rocket wrapped in a ram/scramjet and gives an average Isp of over 1500 secs. It needs to be dual-fuel: UDMH or Methylacetylene (welders MAPP gas is a suitable substitute) for the air breathing modes, LH2/LOX for the high altitude pure rocket mode. I'd also replace the LOX with liquid nitrogen. Yes, its not an oxidizer, but it is half the weight of LOX. As your vehicle is in air breathing mode, you bubble air up through the LN2. This liquifies oxygen in the air and evaporates the liquid nitrogen, so by the time you reach full rocket burning altitude, you have no more LN2, it is now a tank full of LOX.

LOX is normally 80% of a launchers gross lift off weight. Between this technique and the air breathing you can reduce this mass by 5/6ths. By using a denser fuel instead of LH2, you cut your vehicle mass empty by another 30-50% by reducing the tankage.

What does this mean? Lets assume you have a 1 million lb GLOW launcher burning LH2/LOX capable of putting 20,000 lb of useful payload in orbit. Using the above suggestions, you can reduce the vehicle GLOW to under 500,000 lb and automatically dropping the cost per lb to orbit by half.

Nor is, as is claimed by Chris Bell, an SSTO impossible. Putting six space shuttle main engines underneath a shuttle external tank is automatically an SSTO capable of putting 50,000+ lbs into orbit. Did you know a shuttle external tank only costs about a half a million dollars? If you put the six shuttle main engines in a recoverable pod, and left all the external tanks in orbit (to be collected for use in building a space station of proper size for the 21st century), since the engines are the only thing worth saving, you have a cheap to operate reusable SSTO. Carrying 1.6 million lbs of fuel at about $1/lb, you have a total launch cost of $2.15 million (versus the STS' $200-300 million), and a per lb consumables cost of $41/lb. Add in the cost of building and maintaining the engines and engine pods, facilities and launch fees, taxes and other overhead BS, and we are still talking only a few hundred bucks per lb.

Most of the insane $20k/lb cost of the shuttle system is in the 10,000 personnel that NASA employs to rebuild and reinspect every component of the STS for every flight (most of whom are union workers). Learning from the DC-X experience, as well as that of private companies like XCOR, whose EZ-Rocket is capable of multiple flights per day, we can trim assembly and launch staff to a few hundred at most.

Interesting post. I understand it all, but it's more information than I'd be able to put together myself without some serious engineering effort.

Whether horizontal take off is technically "less efficient" than vertical takeoff I think is less important than the ground handling aspects of the two methods.

A winged vehicle with wheels can simply be rolled out to a runway for takeoff at any time. While a vertical launched rocket will require some kind of massive ground handling machinery, and much time and labor to carefully erect it on a launcher.

Those attendant costs of ground handling I believe would erase any efficiencies gained by vertical launches of a vehicle as you describe. Even using the 6 shuttle engines under a throw away tank requires retrieving the engine pod, refurbishing it and transporting it back to the launch point.

There are several "staged" methods of horizontal launch that I believe are also available to improve costs. Staging is expensive with rockets, because retrieval of lower stages is difficult or impossible, and vertical assembly is complex. While staging with winged horizontal aircraft I believe would be much less expensive.

There was a recent test of towing an F-106 into the air to simulate procedures where you could tow a vehicle to altitude to reduce fuel requirements. Also a vehicle was proposed that would use air-to-air refueling to tank with kerosene to power it to Mach 10-12 at 350k+ feet where it would release an upper stage from an internal bay.

Both of these ideas sound like airshow tricks. But they're very easy to do and require no blank-sheet-of-paper engineering. Gliders are towed into the air every day, and literally require a few seconds of ground hookup time. Air-to-air refueling is done daily by the military. Towing a ram/scramjet/rocket to Mach 3+ is very easy to imagine with an XB-70 class aircraft. I know where one of those is sitting in a museum right now.

These techniques aren't within the experience of the rocket men, so I think they're ignored when discussing the issue.

The other serious error is the requirement to boost very large payloads. Yes, very large payloads are occasionally required, and should be done with existing throw-away rocket technology. But small payloads, launched with minimal ground handling required, zero equipment thrown away, and launched almost daily will quickly amortize costs down within reason.

No matter what technique is used, until flights are flown with enough frequency to amortize the development and construction costs across many thousands of flights, then it won't ever be cheap enough for me to afford a flight. Fuel is cheap vs. development, hardware, and labor. Maximum turn arounds mean maximum amortization.

It's an interesting design. It suffers from one main problem, mainly that, given currently forseeable technology, it won't work.

One problem stems from here:

"This engine produces 1,200,000 pounds of thrust, with an exhaust velocity of 30,000 meters per second, from a thermal output of approximately 80 gigawatts. This equates to an Isp of 3060 seconds. Several sources state that a gas core NTR can exceed 5000 seconds Isp, so 3060 is well inside the overall performance envelope. The three turbopumps from the SSME are run at low power levels, and even losing a pump allows the engine to continue running as long as there is no damage to the nuclear core."

The assumptions involved here are too broad and, frankly, wrong. To begin with the closed cycle (i.e. "lightbulb") GCNR variants tend to top out at 20,000 m/s exhaust velocity according to my research. Requiring 30,000 m/s calls for a very high-performance, and possibility of which is questionable.

But this is a mere quibble. Even with 20,000 m/s, performance is vastly increased over chemical boosters and the payload mass would simply need to be toned down. The main problem is in the next part.

"Lets assume this design is able to achieve a thrust to weight ratio of ten to one, so the engine and all of its safety systems, off-line fuel storage, etc, weighs 120,000 pounds. I think we can build this engine easily for 60 tons."

This estimate sounds reasonable, but is, in fact, quite wrong. Basic research would have determined this. The problem is not that closed-cycle GCNRs need be so massive, but rather that they suffer from poor mass flow. You can spit out that hydrogen at very high velocities, but you can't spit out very much of it per unit time. No proposed closed-cycle GCNR, to my knowledge, has achieved a thrust-to-weight ratio greater than one. Ten is completely out of the question. Alas, inspiring as this rocket is, it will never be able to lift itself off the ground.

But all hope is not lost. One possibility is to use an open-cycle GCNR. These benefit from higher exhaust velocities (50,000 m/s is frequently quoted) and higher thrust-to-weight ratios, though I'm not sure if they can achieve ten-to-one. These would probably allow for a rocket similar to the author's specifications to be constructed. The downside is that such a rocket releases its nuclear fuel to the environment at a steady rate. This doesn't particularly bother me, but it would be politically inconvenient. This is the reason that closed-cycle GCNRs were developed in the first place.

Another possibility is to use a solid-core nuclear thermal rocket, which the author briefly touched on but then abandoned in favor of the GCNR. The temperature limitations due to keeping the reactor solid only allow exhaust velocities of around 8000 m/s, but certain solid NTR designs such as the DUMBO type allow thrust-to-weight ratios as high as 70, comparable to a space shuttle main engine, and release no radioactive fuel. This may also allow a similar design.

Keep in mind that I'm only speaking to actual release of the fissionables and their byproducts. Even with a solid-core design, a nuclear rocket of any kind generates intense gamma and neutron radiation when critical. If the rocket carries crew or passengers, some amount of radiation shielding may be necessary. I'm not knowledgable enough about the subject to say how much, particularly given that the hydrogen fuel would absorb neutrons very efficiently, but it needs to be considered.

Alternatively, wait thirty to fifty years for a space elevator to be build, at which point surface-to-orbit rockets will be a moot point.

Anyway, though you'll never find a more adamant advocate of nuclear power and space exploration than me, but this particular design will not work. Sorry.

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