You are not logged in.
Based on an Earth-to-Orbit velocity of 9500 m/s, and a Mass Ratio of 5 (not quite a 20% structural fraction given payload but nevertheless quite high), an exhaust velocity of 600 seconds is needed.
I think that if we can hit 600 s, exponentially averaged over the trajectory, we can do SSTO. Solid Core Nuclear Thermal Rockets can do it, but probably only with Hydrogen and they really aren't safe to use for launch from Earth. Likewise with liquid core. Nuclear lightbulb is possible, but has significant technologies and is a far-off technology regardless.
I'm thinking more about hybrid nuclear-chemical rockets. Something like LANTR, except with the oxidizer introduced in the engine. The idea would be to heat the fuel and oxidizer separately to a high but manageable temperature with a nuclear reactor-- say 1500 C, then introduce them to each other in the engine, allowing the gases in the chamber to reach fantastically high temperatures. This obviously presents an issue in terms of keeping the chamber walls cool, but there are solutions to that, such as circulating non-preheated fuel/oxidizer along the walls (ala Gas Core, Light Bulb) or other methods such as having a layer of char that you don't care about but has an equilibrium thickness.
If done with H2/LOX, this just might be enough to get there. I'm sure there are other ways; Does anyone have any in mind?
-Josh
Offline
Like button can go here
If you have some magical means of convincing people of the safety of nuclear thermal rockets, then it might work. Though I'm very wary of superheated Oxygen...
I've read that Boron mixtures might be able to achieve such an Isp. It was what Mike Lorrey was hoping to use.
Aside from that, and using solar energy to heat the propellent, I don't have any other ideas at the moment. Maybe beamed power and a hybrid chemical-electrical engine? Adding half as much energy again to a Hydrogen/Oxygen rocket might give you 600 seconds.
Use what is abundant and build to last
Offline
Like button can go here
Well, I think I could convince the engineers. 1500 C is on the high side of current reactor temperatures, but is by no means impossible. Especially given that NERVA successfully tested engines at closer to 2500 C. Nuclear technology may not have advanced too much in the time since (though there have been plenty of white papers), but materials technology sure has. While you're right to be skeptical of hot oxygen, it's not an insurmountable issue. Because the Hydrogen and Oxygen will be heated separately, the separate reactors can be designed for oxidizing and reducing environments. That means high temperature oxide ceramics in the Oxygen reactor, and a metal or carbon in the Hydrogen one.
Assuming that the energy components of exhaust velocity add linearly (they don't, but for a first order approximation) H2/LOX provides a bit over half of the energy required to hit 600 s. Assuming a payload fraction of 2%, and a payload of 3 tonnes (enough to launch small satellites, supplies, or conceivably a couple people), the launch mass will be 150 tonnes. Assuming a liftoff T/W of 1.25, the required liftoff thrust is 1.85 MN. At 600 s (The specific impulse will be lower than this at sea level, but so will the input from the nuclear engine), this corresponds to a mass flow rate of a bit over 300 kg/s. Assuming 50% of this energy comes from the nuclear engine, 2.8 GW of nuclear thrust will be needed. That's more or less in line with a nuclear reactor on Earth.
Unfortunately, this seems to suggest specific powers that are fantastically high. If nuclear reactor mass is constrained to 2% of launch mass (3 tonnes), the specific power will be in the region of 1 MW/kg.
How did NERVA do it?
-Josh
Offline
Like button can go here
Nuclear thermal can be made safe. NASA developed NERVA in the late '60s and early '70s, ready for testing in space in 1972. It had 825s Isp at that time. A 1991 study improved Isp to 925s; but that was a study only, no hardware. NERVA used ceramic fuel elements, so it was safe. The problem with NERVA was excessively high reactor mass.
Timberwind was developed by DARPA in 1992. It had dramatically lower reactor mass, but wasn't restartable, and safety was questionable. It wasn't nearly as safe as NERVA. They increased Isp to 1000s, but that was simply by increasing temperature. The problem was temperature was so high that fuel elements melted together; that's why it wasn't restartable. A hybrid between these two seems obvious.
Keeping nuclear safe starts by using uranium instead of plutonium. Uranium oxide is chemically about as toxic as rust. It's only the radiation you have to worry about. Of course Timberwind used uranium metal, which is more toxic. NERVA used ceramic fuel, so it was contained. Fission fragments are the real worry. On the other hand, plutonium is so poisonous that if you got a piece the size of a grain of sand in a cut, you're dead. So one issue is educate the public: only plutonium has all those nasty effects they're worried about, not uranium.
If the goal SSTO, then consider a jet engine. The Saturn V used RP1/LOX for its first stage, LH2/LOX for its second. The Shuttle used solid rockets for boost, and LH2/LOX sustainer. So why not use a jet engine for take-off and powered landing, and solid core NTR with LH2 for the push into orbit. There are a couple options. Rocket-Based Combined Cycle (RBCC) is an engine that starts as conventional RAM jet, transitions to SCRAM jet, then air augmented rocket, finally becoming a conventional rocket engine with LOX. The term "air augmented rocket" could be called "LOX augmented jet engine", depending on your point of view. Is the glass half full, or half empty? This has been done in a test stand using LH2 only. Of course integrating a SCRAM jet with an air frame is non-trivial. One proposed project was to do this with a turbine jet engine, which would transition to RAM jet, bypassing the compressor and turbine. That proposal also hoped to use kerosine based jet propellant. The X-program was cancelled; if the Air Force proceeded anyway then it's become secret. But NASA is funding research into SCRAM jet drones/missiles, the program goal is to achieve mach 20. If they get it to work, that would make a great take-off/landing engine. I can't call it "first stage" if nothing drops off.
A more radical option I've mentioned before is a nuclear thermal jet engine. The Air Force did some research on this in the 1950s. Some nuclear engineers said if it were done today they could probably get it to work. Achieving most of your orbital speed without any carried propellant at all would be of great benefit. To be useful it would have to be used for ground launch. And achieve somewhere around mach 20 at high altitude. If it could only achieve mach 10, it would be questionable whether it's worth it.
Last edited by RobertDyck (2013-09-15 16:29:12)
Offline
Like button can go here
Indeed, it seems that the NERVA engine's T/W, as developed, was just 3. I've heard that the DUMBO proposal would be able to raise this to 10 or so, but that's still not really high enough for the engine on an SSTO rocket.
I wouldn't go so far as to call NTR "safe". First of all, of course, one has never been flown. Beyond that, accidents do happen. In a chemical rocket, the rocket explodes. (Who's seen the video of the Soviet N1 blowing up, or more recently the Proton?) In an NTR, you have a massive nuclear reactor with hundreds of kilograms, if not tonnes, of highly radioactive Uranium and/or Plutonium everywhere. Depending how close to the launch site this happens, and which way the wind is blowing, that's an excessive danger, and even if it happens far from the launch site the environmental effects are still pretty bad.
I tend to agree with Nyrath when he says (on the atomic rocket website) that nuclear fission is safe as long as you treat it with the respect that it deserves. Shooting multigigawatt reactors off at high temperatures and hypersonic speeds probably doesn't qualify. The hybrid that I suggested is easier to build in a safer way because it's at a lower temperature and can be separated from the rest of the engine somewhat, but still probably falls under this heading.
In terms of airbreathing rockets, Scramjets are hard. There are also significant questions, in my mind, about whether it's really possible to design a craft that could withstand prolonged periods of high velocity travel in what GW Johnson likes to call the "sensible air". Likewise, it's very difficult (perhaps impossible?) to build a common engine that could handle airbreathing combustion at that range of velocities.
I do like the idea of a jet engine; Even if it only has the effect of shaving 700 m/s (Mach 2) off the delta-V of the rocket engine and gets you up to a region of low pressure, the mass ratio for the rocket stage (assuming an Isp of 450 s) need only be 7. We're not quite there, of course, but that's an improvement. To get down to a Mass Ratio of 5, the delta-V would have to be decreased further by about 1.6 km/s (for a total non-rocket delta-V of 2.3 km/s) or the Isp of the rocket fuel increased to about 550 s. In energy terms, this is a pretty big deal. To be clear, I assumed for the purposes of calculation that the Isp of jet engines is infinite, which at an effective Isp of 3000 s and a delta-V of 700 m/s it nearly is.
The downside to jet engines is their T/W. According to Wikipedia they generally will not exceed 8.
While we're talking about T/W of engines, apparently the Merlin 1D has the best T/W of any engine ever fired, at about 160.
-Josh
Offline
Like button can go here
NERVA "did it" with limited-life and liquid hydrogen at a fully-ionized MW of 1. If we were to reprise NERVA today, I'd think Isp 900 s at engine T/W 4 would be quite feasible for a very first article, again with LH2. But, it's LH2. Problem!
I'd certainly do that again for a first article, but I'd also start a parallel effort to solve the chemistry and materials problems with a "water NERVA", as in the sci-fi of the 1940’s and 1950’s. Performance might be somewhere near the 600 s Isp we are discussing here (not close to 1000 s), again at engine T/W 4 in a first article.
That's actually getting very close to what Josh started this thread about. Follow-on articles would look more like Dumbo (metallic instead of ceramic reactor fuel containment) than Timberwind (a fluidized-bed particle-layer reactor), at T/W approaching 10. If not SSTO, then that gets you a very practical TSTO with a minimal throwaway 2nd stage, and perhaps a recoverable winged 1st stage working VTO and HL.
The main attractive thing about a water NERVA is its rework into a space engine instead of a launch device to LEO. Avoids all the safety-of-flight concerns regarding damage to the public.
Plain water is very easily transported as icebergs, and is also almost impervious to meteoroid impact damage that way. Water-as-ice is located on a variety of bodies, including Mars, and might well be fairly easily mined, depending upon the exact nature of the deposits (which also points out how important prospecting/ground truth really is, as we explore with men).
Separation of solid contaminants from liquid water is SO very easy, even if you have to use spin-derived artificial gravity. My best guess is that any solution to the chemical/material challenge of a water NERVA could easily handle CH4, NH3, and similar liquid contaminants in the water, so purity should not be much of an issue, although salts might be. Be nice if our NERVA could handle salt water, wouldn’t it? Think 2nd-generation device.
How could one possibly "beat" a high-Isp/high-thrust engine (with VERY easily-obtained propellant throughout this solar system)? It's a dream-come-true!
The follow-on is something that deserves very heavy-duty development, too. Consider an open-cycle gas-core water-propellant nuclear thermal engine as a deep space engine. If I had to hazard a guess, I'd say engine T/W could be near 30 at Isp near 2000 s. The real limitation is that regenerative cooling won't be enough for the shell and bell. But, there is no core or core-containment to damage or melt, it's already plasma. (This might also mitigate some of the safety-of-flight issues for nuclear launch, too.)
All of that is fine for relatively-small vehicles. But, for colony-planting vehicles, nothing (and I do mean nothing!!!) beats the nuclear explosion drive. It works best in gigantic sizes (10^4+ tons), so there's little point to using it in small vehicles, as NASA "tried" to do in 1965, looking at it for the manned Mars mission then planned for 1983. They went with NERVA instead, and chemical as its backup. (Lost-water-under-the-bridge, since all of that was cancelled in 1972, in the middle of the Apollo landings.)
Beyond that, we're looking for Star Trek-style impulse engines and warp drive. The physics ain't there yet for them, much less the technology.
As for “combined-cycle anything” in chemical propulsion, the geometry changes have always been the bugaboo. Anything we can practically do for changing geometry very badly (usually fatally) compromises the performance of the individual cycles.
For example, subsonic combustion ramjet usually features a constant area combustion chamber contracting toward a throat by a throat/chamber area ratio 0.65 max, and not much less. The exit bell falls in the 1.5-2 area ratio range. Period. That’s what works. Anything far from that essentially loses all the performance potential. I know, I used to design ramjet missiles for a living.
Scramjet (supersonic-combustion ramjet) on the other hand features (at worst) a constant area chamber, and usually gently-expanding at around 5 degrees half angle. There is no throat contraction at all, but locating the final expansion bell axially can make-or-break obtaining a burn at all. This technology is still very far from being ready-for-prime-time, the recent X-51 flights notwithstanding. There’s no good way to reconcile these conflicting geometries except by one-shot/throwaway ejected components, and even that is very most certainly not a trivial exercise, or a “sure thing”.
And, we have not addressed inlet geometry incompatibilities between ramjet and scramjet at all. They are huge, more especially in the internal ducting lines, surprisingly enough. The external compression features are actually quite similar, which is terribly misleading. Failure to get this right causes as many violent explosions in scramjet test articles as does too-low a scramjet takeover Mach. It is quite catastrophic, and (so far) quite incompatible.
Integrating ramjet-or-scramjet with rocket is even worse. Most rockets have a very large area ratio contraction from chamber to throat: on the order of 10+. And the exit bell area expansion ratios exceed 10, often by a very, very large margin. I know of no variable-geometry techniques to accomplish this kind of geometry change, except one-shot/throwaway ejectable insert items.
You can go from rocket to ramjet that way, but you absolutely cannot go back to rocket. How are you going to change back to rocket in a combined-cycle engine, if your trade studies say that you want to do that? I don’t think anybody on Earth has a practical answer to that, excepting maybe the Skylon folks with their SABRE engine, and I am very definitely not even sure of that! (I hope they do, but I am most definitely not going to bet the farm on it.)
And that’s why I think parallel-burn options for the differing engine types are way-to-hell-and-gone far superior to any combined-cycle proposals I have ever heard of. I have seen many of those for the last 4.5 decades. None has ever led anywhere, before. Not a good track record.
But parallel-burn works, both ways. Try that. We can do it right now. All-existing technologies. Not trivial, but very definitely do-able.
GW
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
Offline
Like button can go here
Actually, I would call NERVA "safe". It used ceramic fuel elements. The uranium wasn't just encased in ceramic, it was evenly mixed with ceramic. The elements were solid. So in case of an explosion, the uranium would not be "dust" blown in the wind. Rather it would be solid pieces, fallen in the strewn field. And they went further: NERVA was designed to be launched cold, the reactor never turned on. That means all uranium, with no fission fragments at all. So those pieces could be picked up with nothing more than loose plastic gloves that come with oven cleaner.
From July 1987 through July 1990, I lived in Toronto. One day I watched a show on TV Ontario, their equivalent to PBS. This show was about the CanDU reactor. It showed workers filling fuel rods. They wore white lab coats, plastic shower caps, simple lexan protective eye wear, and the same loose plastic gloves that come with oven cleaner. That's where I got this. The workers were shown stuffing yellow cake powder into stainless steel tubes with their fingers. A bundle of these tubes were the fuel for a CanDU reactor. So the uranium fuel elements were assembled by hand. They are that safe before they go into the reactor. However, when they come out it's comletely different. They have solid concrete walls several feet thick separating spent fuel elements from any humans. A robot carries the fuel to a water storage facillity, which looks like a swimming pool. That pool has a bridge crane to move fuel elements. So big difference between uranium vs fission fragments.
However, I did suggest ground launch. That would make the reactor hot. Still, although spent fuel would be highly radioactive, it still wouldn't blow in wind. The ceramic was designed to keep individual pieces intact. Even if one broke, it just exposes another face. Nothing spills out.
Last edited by RobertDyck (2013-09-16 09:07:15)
Offline
Like button can go here
For use strictly in space, yea, open-cycle gas-core NTR. But for ground launch, or anything in the atmosphere, I argue for solid core just for safety. And anything launched from the Earth has access to all our industry. Liquid hydrogen requires special handling, and it's light but the tank is extra large. Still, Isp and total launch mass are improved so much that you just have to.
Last edited by RobertDyck (2013-09-15 17:16:27)
Offline
Like button can go here
Boeing claimed they got an RBCC engine to work in a test stand. The diagram showed how they dealt with geometry. Pardon me if I use the wrong terms. Rather than circular cross section, they used rectangular. Just as SuperHornet or F-15 Eagle has a rectangular cross-section inlet, but they did so for the entire flow path. To change geometry, the had floor and ceiling sections tilt and contract/expand. And multiple fuel injectors; as the engine flow speed changed, they switched fuel injectors. That moved the position forward or back. And to change to rocket mode, they contracted the inlet to completely seal it off, and contracted a "hump" part way down to create the rocket "throat". They claimed it worked.
When the X-prize started, I sent an email to Boeing asking for more information. My idea was to use their RBCC engine for an entry. But their email response was to ask who I was, and they took their website down. :-( Was it all vapourware? A scam to get research money from the military?
Offline
Like button can go here
Robertdyck-- I think we're approaching the safety issue in different ways. I don't disagree that, in the normal course of operations, an NTR is "safe" insofar as normal operation does not spew radioactivity everywhere like open cycle does, and it doesn't spew out toxic chemicals. While the exhaust may contain an elevated amount of deuterium, one could hardly call that a significant issue. What it does do, however, is produce a lot of radioactivity at a very high rate while it's firing. I am aware that aside from its tendency to burn vigorously in the elemental form, Uranium is not particularly poisonous. However, I would like to add that the Uranium used in CANDU reactors is mostly natural Uranium, meaning that it contains just .7% of the good stuff, U-235. Even U-235, before it's been fissioned, is not particularly bad from a radioactivity sense (I believe its half life is many millions of years, perhaps a billion?)
Coming out of a reactor of any kind is a different story. The materials out of which the engine are built will become highly radioactive; And I don't care what ceramic the Uranium is dispersed in, I bet you anything that it won't survive impact with ground or water at 5 km/s. And forget about landing in an inhabited area-- that engine is going to be far too hot, from a radioactivity standpoint. For that matter, forget about reusability entirely, because you'd have to go through the understandably expensive and time consuming procedure of replacing the intensely radioactive nuclear fuel at each launch. A 1% failure rate is pretty good for disposable rockets, I'd be surprised if adding the reentry step decreased this for any reason. The idea of having one of these crash anywhere near an inhabited area is rather terrifying.
-Josh
Offline
Like button can go here
An airliner crashing into an office tower is terrifying. I don't think crashing a nuclear powered rocket would be any different. Just stay away from the highly radioactive fuel elements lying on the ground. You should stay away from debris of a crash anyway.
As for reusability: the reactor could not be refuelled, it would have to be completely dismantled to refuel. But a certain minimum critical mass is required for the reactor to function. That is more than fuel necessary for a single launch. So the reactor could be used for multiple launches without refuelling it's uranium. Modern neutron reflectors could reduce the critical mass. And using U-233 instead of U-235 would reduce the critical mass further. Yes, I'm all for carrying the minimum amount of uranium. But even with all that, it would have enough for multiple launches without refuelling. It would only need to refill propellant, manoeuvring thruster fuel, etc.
Offline
Like button can go here
Non-nuclear alternatives:
The X-43C was the X-program to develop a turbine based combine cycle engine. It was supposed to use kerosene based jet propellant. Not RP1, I said jet propellant. Don't know which grade, but I don't think they pinned it down. The program goal was to build an engine that could take off and land from a runway on its own power, and accelerate to mach 6. The military was very interested for obvious reasons. I have visions of mach 6 jet fighters. But NASA was interested as a component for a future space shuttle.
NASA had another research program. This was to produce an RBCC engine that was strictly RAM jet. It would be thrown into the air with a linear magnetic catapult. It would use LH2 fuel, and fly to high altitude and high speed. Program goal was mach 10 minimum, but they were hoping for mach 17. The RBCC engine would transition to LOX/LH2 rocket for the final push to orbit. I'm sceptical of any catapult that throws an aircraft into the air at mach 1; the minimum speed to ignite a conventional RAM jet. You don't want to be close to the ground when the sonic boom hits. Ground reflection and all that.
If you want separate engines, could you do this with a jet propellant engine for take-off/landing, accelerate to mach 6. Then switch to pure SCRAM jet using LH2 for acceleration from mach 5 or 6 to mach 20. Then switch to separate LOX/LH2 rocket for the final push to orbit. That means 3 engines. It also means you could use the JP fuelled jet engine for powered landing. That would enable holding patterns and fly-around, so landing at a commercial airport.
Last edited by RobertDyck (2013-09-16 09:33:42)
Offline
Like button can go here
You want to fly in sensible air at Mach 20? Seriously?
Use what is abundant and build to last
Offline
Like button can go here
You want to fly in sensible air at Mach 20? Seriously?
NASA has given a contract to ATK to explore SCRAM jet technology, with the program goal of achieving mach 20. That's why I mentioned that number. Seriously.
Offline
Like button can go here
Perhaps the confusion here is with regards to what constitutes "sensible air"? GW would obviously be the one best suited to say, but I think he usually sets the limit at around 60,000-80,000 ft altitude, or 18-25 km. While still a very tough engineering challenge, this is not nearly as bad as going mach 20 in the breathable air. At these altitudes the pressure is between about 4 and 10 kPa, depending where exactly in the range you are. Whether going almost 7 km/s without destroying your craft is possible is not a question that I am able to answer; But a scramjet should be able to operate even higher than this, given its faster speed.
I have serious doubts about the feasibility of a scramjet, especially for reusability applications, but it's not unreasonable to suggest in a thread where the opening post was about a hybrid fission-chemical rocket.
-Josh
Offline
Like button can go here
If everyone will tolerate my returning to hybrid chemical rockets for just a bit, I really do think it's an idea with fascinating potential. Matter moving at 6 km/s has a kinetic energy of 18 MJ/kg; matter moving at 4.5 km/s has a kinetic energy of about 10 MJ/kg. Hydrogen/Oxygen fuel has a lower heating value of 13.3 MJ/kg, giving an engine efficiency of 80% (!!!). Because the engines burn at an extremely high temperature, this is still below the Carnot (physical limit of) efficiency, which at SSME combustion temperatures is 91%.
I will assume that increasing the exhaust velocity has no effect on efficiency. This means that we need a fuel with an energy content of 24 MJ/kg, or a fuel that has had energy totaling 24 MJ/kg added to it. On top of the energy contained in H2/LOX, this is 11 MJ. Interestingly, on top of the energy contained in Aluminium/Oxygen, it's just 9.5 MJ, although this has some issues with aluminium oxide being, ya know, a solid.
On the other hand, there is another way to go about this: Increase your effective Isp without increasing the energy used significantly. How do you do this? It's actually rather simpler than it sounds: Airbreathers. I suspect that if you don't desire to use the oxygen in the air as an oxidizer, but rather simply as reaction mass, it becomes much simpler to build a workable engine. The principle, of course, is simple: Energy is proportional to mass and velocity squared. Momentum is proportional to mass and velocity. Therefore, by keeping the amount of energy constant but decreasing the velocity, you can use significantly more mass. If you have a source of free mass outside of what you need to carry (I like to call this "Earth's Atmosphere") this is a viable option if you can get the engineering to work.
Assuming a reduction in efficiency to 75% due to anything with an air intake being less efficient than a rocket by nature, at a speed of zero one would need a mass ratio of 1.8, which is to say that the for every one kilogram of rocket propellant, you will need to intake .8 kg of air. As your speed rises, this does also. The general relation is:
R=Veff/(2E/Veff-4Vc)
Where R is the ratio of reaction mass including air to the reaction mass excluding air, Veff is the effective exhaust velocity (in this case, 6000 m/s), E is the specific energy of the propellant times the efficiency, in this case 10 MJ/kg, and Vc is the velocity of the craft. The ratios are as follows:
0 km/s 1.80
1 km/s 4.50
2 km/s N/A
3 km/s N/A
4 km/s N/A
5 km/s N/A
6 km/s N/A
7 km/s N/A
8 km/s N/AMy equation suggests that above 1.7 km/s (M5), it becomes impossible to transfer enough energy to the remass to actually get a full 600 s effective Isp out of it. At Mach 5, a ramjet (according to Wikipedia) gets an effective specific impulse of up to ~1500 s. I'd like to hear actual numbers on that from GW, but assuming that the actual exhaust velocity remains constant, while a ramjet simply reduces the onboard fuel storage (the ratio of fuel to oxidizer is 1:2.3 in a kerolox rocket), I would expect the maximum Isp to be around 1900 s based on the calculations done here. On the other hand, if they burn fuel rich that could easily be brought down; increased friction, decreased efficiency, less pure fuel with a lower energy density, worse performance of oxygen/nitrogen/CO2 as propellants compared to water vapor, lower burn temperature of hydrocarbons in air relative to Hydrogen in Oxygen resulting in a lower engine efficiency*, and any number of other factors, as well as the approximate nature of the graph I'm reading this from on Wikipedia, could easily explain the difference between my ideal situation and the actual value.
*This is calculable: Relative to the 3500 K flame in the SSME, Kerosene in air burns at 2350 K. This results in a decrease in the Carnot efficiency from 91% to 87%, a 3.5% reduction, which explains about 1/8 of the difference between my calculated and the (approximate) actual value. I have neglected the effects of compression heating for the purposes of this calculation.
-Josh
Offline
Like button can go here
Looking at the post that started this thread, in the hybrid chemical-nuclear engine we are talking about adding oxidizer and burning chemically, after heating the hydrogen through a NERVA-like scheme.
In that kind of scenario, the temperatures will go high enough to ionize the propulsion stream quite significantly. The more chemical combustion you add, the more ionization you get. That's the same problem that limits plain ramjet combustion to Mach 6 airspeeds: combustion temperatures getting into the big-time ionization range when they exceed about 4500 F (2480 C). All chemical rocket engines already suffer a little from this effect, since they burn with chamber temperatures in the 5000-6000 F range (2760-3320C).
Why ionization is a problem: the nozzle converts internal energy to kinetic energy, but it does not recover anything from that portion of the applied energy that went into ionization. The recombination contributes nothing to exhaust velocity.
Somewhere, I think I saw a proposal to afterburn a NERVA-like device with injected oxygen. But the injection point was somewhere in the nozzle bell, to avoid the ionization problem.
It's something to worry about, anyway.
Sensible air is really a concept for airbreathers and folks calculating drag forces. It's a fuzzy limit, depends on how large the forces are, in which you are interested, compared to the other forces in your problem. Ramjet, scramjet, and turbine thrust forces tend to get rather small compared to vehicle weights, somewhere in the 60-80,000 ft range. Vehicles propelled only by those kinds of devices tend toward 0.01 gee (or less) accelerations at around that altitude, instead of the 0.1 gee+ we'd like to see.
GW
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
Offline
Like button can go here
GW-
The reason why that energy isn't useful is that it's released as radiant energy, instead of as kinetic energy of the molecules, right? In that case, could the energy be retained within the exhaust stream by utilizing fuels that have some solid products that would be opaque to the radiation released thereby, and thus keep the energy within the combustion stream?
-Josh
Offline
Like button can go here
Well, now that's an interesting idea. I hadn't really thought about anything quite like that. The radiant energy isn't released until the recombination starts, but once it is, a solid content at a concentration just high enough to be optically-dense just might do the trick to absorb it.
Then you are faced with the heat transfer problem of getting it from the solid absorber into the gas. The solid particulate will really have to be extremely fine to speed up that process, because about the only parameter you have to control heat transfer rate is surface area. I'd suggest fine soot.
GW
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
Offline
Like button can go here
Will recombination occur within the engine? Much less absorption and transfer?
By the way, gas will absorb radiant energy. Depending on the energy frequency/wavelength, and which gas, and gas density. If recombination and radiation occurs in the bell near the throat, then the solution could be a simple fuel additive to produce a combustion gas that absorbs it.
Offline
Like button can go here
I would expect the emission to be in the ultraviolet, but that's just a guess. It's my understanding that most gases are not that optically thick, on a per-mass basis, in any spectrum, though this isn't something I consider myself to know very much about. Certainly on the distance scales we're interested in Compton scattering isn't going to be much help. I was thinking more along the lines of Aluminium being added, because of its high combustion energy, but I suppose Carbon (if you could keep it in a true carbon form coexisting with CO2, as opposed to them reacting to CO (I believe this is thermodynamically favorable). I don't know how much control we would have over the particle size, but I would expect that given the (probably turbulent, as opposed to laminar?) flow conditions within the combustion chamber and engine bell, the typical particle size would be quite small.
This "thermal rocket effect" as I like to call it, is observed in the real world, by the way: It's presumably a vital part of the function of any solid rocket engine. They're all Aluminium fueled, after all, and Aluminium Oxide is sure as heck not a gas at Solid Rocket operating temperatures.
It presumably necessitates a longer nozzle. Aerospike, perhaps?
-Josh
Offline
Like button can go here
Adding a dust to absorb radiant energy, with the hope of heat transfer to a gas? All quick enough to cause gas expansion before it leaves the rocket exhaust nozzle? No way. Either go with gas, or forget it. Besides, do you realize how much errosion you would get from dust in a rocket engine?
But I'm going to repeat the first question. Where does recombination occur? Is it even within the engine at all?
Offline
Like button can go here
Recombination does not start until you reduce internal energy, either by heat withdrawal (for some other purpose), or by the start of expansion. The recombination that happens in liquid and solid rockets occurs in the exit bell, usually upstream of the throat (effectively-frozen chemistry in the supersonic part). Most of the thermochemical codes that compute Isp or c* for you take this into account. Going at it pencil-and-paper from just theoretical energy content, you fail to account for this.
If I was going to seed the flow of an LH2 NERVA to make it more opaque, I'd inject carbon monoxide just downstream of the core. I think the hydrogen would steal the oxygen from the monoxide, leaving at worst colloidal-size soot in the flow, maybe molecular size. The effective emissivity of a solid like that is usually between 0.1 to 1 if optically dense, while the effective emissivity of straight furnace gases is usually in the 0.01 to 0.1 range, at most. Could be lower.
How all that plays into a hybrid chemical-nuclear device, I am unsure at best. But, it sounds like multiple "treatment" stages, some before the throat, some after.
Aluminum in solid rockets seems to be an additional energy source in the chamber, and a loss in the nozzle flow process. In the solids, the reduced-smoke composites had little or no metal, and flame temperatures near 4500-5000 F. With 20% aluminum, flame temperatures were closer to 5500+F, which means it added almost 1000 F. The product is liquid/solid aluminum oxide, which contributes nothing to the hot gas expansion process. You have to remember that these formulations are solids-limited (thixotropic mix "viscosity"-limited), not oxidizer-limited. They are nearly always badly-underoxidized. The aluminum replaces some of the AP.
The droplets/particles are relatively large, and there is no time for any effective heat transfer. What you get at the exit is a high speed gas jet filled with 5500+ F sand blasting "grit" that lags a bit in velocity behind the gas. It's really tough on O-ring seals, as they found out back in 1986 with Challenger's SRB joint.
If you add to that, the fact that O-ring failures are nearly always concentrated at a single point, then you can understand why old hands like me see no point to using more than one O-ring in a solid rocket motor joint. The 3-ring joint they finally came up with was even less reliable than the 2-ring joint they started with. Only reason it never failed again was that they never flew it 29 F cold again.
GW
Last edited by GW Johnson (2013-09-18 08:46:18)
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
Offline
Like button can go here
In any case, given the Thrust-to-weight considerations, I don't think a chemical-nuclear hybrid rocket is the best idea for SSTO. Regardless, I don't think that putting CO in the exhaust stream would result in C+H2O, but rather just remain (mostly) unreacted. Given that the entropy is decreasing (2 molecules of gas are becoming one of gas and one of solid) and the reaction absorbs energy (C=O is a very strong bond) I'd imagine that this would be relatively stable configuration. Methane might be better.
Regarding heat transfer, having not done the equations at this time it may well be justified to suggest that it's not workable.
-Josh
Offline
Like button can go here
Josh:
Don't give up yet, but my numbers say you want closer to 700-800 sec Isp than 600. I did a trade study and posted it over at "exrocketman". It's the 9-24-13 posting.
GW
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
Offline
Like button can go here