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A number of years ago, the President of France said many countries have technology to stop a missile. He said if Iran tried to launch a ballistic missile with a nuclear warhead at Israel, France would stop the missile before it got 10 metres above the ground. And France would then nuke Iran. I had posted on this forum that someone should say that. Yea! But then NATO allies got him to retract that statement.
Israel has nukes. Arab countries are going to panic until they have nukes as a counter threat.
Seems to me the Arab countries are more afraid of Iran, than of Israel. There is a difference between a responsible nation and an irresponsible one. Iran is an irresponsible nation, it employs suicide bombers, and uses terror tactics in pursuit of its objectives, it seeks to dominate the middle east and is therefore a threat to all establishment powers, and I don't think you want an Islamic Theocracy ruling the World do you? But that is Iran's objective, and such an objective is incompatible with World peace, so letting Iran have nuclear weapons is threat to the survival of the human race, just as surely as an asteroid impact is.
But they can't use the Nukes. If Israel does, then Arab nukes would be launched. If any Arab country nukes Israel, then NATO would sanction turning that Arab country into a radioactive hole in the ground. If not French nukes, then the fact France feels that would would give permission for the US to do it. And Israel would definitely launch their unless they were disabled. MAD: Mutual Assured Destruction.
All the more reason to build an atomic bomb spaceship to get away from all that religious lunacy that threatens to destroy the human race!
In one war, Arab countries invaded Israel and it looked like they were going to win. Israel ordered deployment of their nuke. A bomber took off for an Arab target with a nuclear bomb. However, a small Israeli force managed to halt progress of a major Arab attack force several times their size, and hold them long enough for Israel's main force to arrive. So the attack was thwarted. Israel recalled their bomber, it turned around and returned to base without dropping the bomb. This demonstrates the danger of just one side having the bomb.
If one side had the nuclear bomb, there wouldn't be as many casualities, there wouldn't be mutually assured destruction, Israel would use only as many nuclear bombs as it takes to defeat the enemy that is invading it, and that would be much fewer than an even-handed two way volley of nukes that you advocate. the problem is there are a lot of Islamic fanatics in the middle east that wouldn't be deterred by Mutually assured destruction, so Islam + Nuclear weapons = A threat to the survival of the human race!
Perhaps a nuclear deterrent would stop wars, stabilize the Middle East.
And perhaps it won't, Muslim fanatics don't seem to have the same survival instinct that the rest of the human race does, there are just too many suicide bomber running around blowing themselves up, and allowing them to have nukes is a very bad idea! But if you aren't afraid of that, then I think we should go full bore with that atomic bomb spaceship to save what members o the human race we can from that Islamic Nuclear threat!
I would prefer neither side have nuclear weapons, but I don't get a say. Just trying to figure out what's going on. Once Israel got nuclear weapons, it was a matter of time before at least one Arab country would too.
Why is that, why are Arabs naturally expected to hate Jews?
My understanding is the Iranian deal would let Iran have nuclear reactors for commercial power only, not weapons. We'll see how that works out.
And if it doesn't work out, then we lose our cities and millions of people who live and work in them, are you willing to take that risk? If you are, then you wouldn't mind an atomic bomb spaceship, because the risks of that are much less than a nuclear armed Iran.
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Can we have a discussion about what methods of nuclear propulsion would be useful within Earth's atmosphere (i.e. advantages and disadvantages) instead of talking about the politics of certain groups of people who have or don't have nuclear weapons?
This is a technical discussion forum and the subject of your original post pertains to using nuclear pulse detonation propulsion for super heavy lift required for colonization of Mars. We do have a forum for discussion of political agendas, if that's your cup of tea. Personally, I come here for the interesting technical discussions. I can turn on the TV to virtually any channel and get someone's opinion on politics.
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Can we have a discussion about what methods of nuclear propulsion would be useful within Earth's atmosphere (i.e. advantages and disadvantages) instead of talking about the politics of certain groups of people who have or don't have nuclear weapons?
This is a technical discussion forum and the subject of your original post pertains to using nuclear pulse detonation propulsion for super heavy lift required for colonization of Mars. We do have a forum for discussion of political agendas, if that's your cup of tea. Personally, I come here for the interesting technical discussions. I can turn on the TV to virtually any channel and get someone's opinion on politics.
Wasn't my idea to go political, I was Roberts. I'm just comparing the relative risks between an atomic bomb spaceship, and evil people that we are letting develop nuclear weapons, and with a bit of sarcasm, I'm stating that our fear of nuclear weapons must not be that great since we are letting dangerous people build them. It was Robert that started talking about Israelis and politics.
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Is there any way to tone down the political nuclear as we know that its time for all nations to grow up with this level of power as it does have great potential to those that can control it for the future of the world's peoples.
With regards to the material we have a choice to go small for orbit an beyound and for surface use on other planets as a developmentstep. I think until we can show no meltdowns are possible or some other sever happenings that could have a dier results to man in any surounding area where it is used we will always have resistance to making use of nuclear.
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Maybe its time to show that we can linkup multiple RTG units on mars to allow for base building to start before we do go. This would be a great task for tele robotic use....
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Timberwind produced 1000 second Isp. The last version of NERVA produced 925 seconds. But Timberwind had a problem of producing hot spots in the pebble bed, which agglomerated the pebbles together. That made it non-restartable, single use. NERVA took a different approach, but the reactor was heavy.
Timberwind 45 - rocket stage
Status: Development ended 1992. Thrust: 441.00 kN (99,140 lbf). Gross mass: 28,000 kg (61,000 lb). Unfuelled mass: 7,500 kg (16,500 lb). Specific impulse: 1,000 s. Specific impulse sea level: 890 s. Burn time: 449 s. Height: 23.87 m (78.31 ft). Diameter: 4.25 m (13.94 ft). Span: 4.25 m (13.94 ft).
Thrust (sl): 392.800 kN (88,305 lbf). Thrust (sl): 40,050 kgf. Engine: 1,500 kg (3,300 lb). Thrust to Weight Ratio: 30.
NERVA NTR - engine
Status: Study 1991. Thrust: 333.40 kN (74,951 lbf). Unfuelled mass: 8,500 kg (18,700 lb). Specific impulse: 925 s. Diameter: 5.00 m (16.40 ft).
Engine: 8,500 kg (18,700 lb). Area Ratio: 500. Thrust to Weight Ratio: 4. Coefficient of Thrust sea level: 0.
Notice Timberwind 45 has greater thrust and lower engine mass. Could we design a nuclear thermal engine as light as Timberwind, yet doesn't melt itself?
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I recall a topic on pebble reactors here so will find its link in a bit to add to this post.
searched pebble reactor http://newmars.com/forums/search.php?se … =951986922
While the rocket engine research was stopped long ago the actual reactor however was not.
early version https://en.wikipedia.org/wiki/AVR_reactor
second generation https://en.wikipedia.org/wiki/Pebble-bed_reactor
latest variation https://en.wikipedia.org/wiki/Pebble_be … ar_reactor
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Is there any way to tone down the political nuclear as we know that its time for all nations to grow up with this level of power as it does have great potential to those that can control it for the future of the world's peoples.
With regards to the material we have a choice to go small for orbit an beyound and for surface use on other planets as a developmentstep. I think until we can show no meltdowns are possible or some other sever happenings that could have a dier results to man in any surounding area where it is used we will always have resistance to making use of nuclear.
I just want equal time for Orion, that's all! If we give $400 million to a nation that wants to build nuclear weapons, we should also spend $400 million to build an Orion Spaceship, I'd say fair is fair, if people that want to kill me are getting $400 million from my government to build nuclear weapons, I'd say we should spend at least that amount on building a nuclear Orion.
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Timberwind produced 1000 second Isp. The last version of NERVA produced 925 seconds. But Timberwind had a problem of producing hot spots in the pebble bed, which agglomerated the pebbles together. That made it non-restartable, single use. NERVA took a different approach, but the reactor was heavy.
Timberwind 45 - rocket stage
Status: Development ended 1992. Thrust: 441.00 kN (99,140 lbf). Gross mass: 28,000 kg (61,000 lb). Unfuelled mass: 7,500 kg (16,500 lb). Specific impulse: 1,000 s. Specific impulse sea level: 890 s. Burn time: 449 s. Height: 23.87 m (78.31 ft). Diameter: 4.25 m (13.94 ft). Span: 4.25 m (13.94 ft).
Thrust (sl): 392.800 kN (88,305 lbf). Thrust (sl): 40,050 kgf. Engine: 1,500 kg (3,300 lb). Thrust to Weight Ratio: 30.
NERVA NTR - engine
Status: Study 1991. Thrust: 333.40 kN (74,951 lbf). Unfuelled mass: 8,500 kg (18,700 lb). Specific impulse: 925 s. Diameter: 5.00 m (16.40 ft).
Engine: 8,500 kg (18,700 lb). Area Ratio: 500. Thrust to Weight Ratio: 4. Coefficient of Thrust sea level: 0.
Notice Timberwind 45 has greater thrust and lower engine mass. Could we design a nuclear thermal engine as light as Timberwind, yet doesn't melt itself?
Probably yes. In a power reactor operating at 100% for a long period, decay heat is 8% of reactor power at shutdown, but decays to ~1% within about 60 seconds. For a nuclear rocket engine the decay heat will die off much more rapidly, as there will be relatively few long-lived fission products within the core. So within minutes, decay heat will have declined to <1% of shutdown power. In space, you would need to dump that heat by radiation, so the outer skin of your upper stage will need to be covered by a radiator of sorts. You could work out accurately what the decay heat would be as a function of time, but would likely need a spread sheet that integrated the decay profiles of all of the fission products and their daughter products. Not actually a technically difficult thing to do, just very time consuming.
Use the steffan-boltzman equation to work out how much heat a reasonably sized radiator could dump at an acceptable operating temperature. My quick back of the envelope calc indicated that a cylindrical radiator with the diameter and length of the Timberwind upper stage, operating at 1000K, would lose about 18MW of heat by radiation. That's 0.32% of the 5000MW operating power. So its probably doable. A precise decay profile will tell you how long you need to maintain residual propellant flow through the core after shutdown. You would trip the reactor at something like 99% of orbital velocity and allow the decay heat to provide the final boost into orbit.
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NTR solid core provides around 1000 s Isp at around engine T/W near 5-15. That's about the range this technology can produce. They had to work very hard to reduce the hard radiation in the exhaust stream. If I had to guess, the pebble-bed "welding" experienced with the Timberwind concept led to hard radiation in the exhaust stream again. NERVA-type NTR is a ready-to-employ technology, Timberwind not so very much, it still needs some work. This technology is suitable as an upper stage in something larger, or as a very limited single stage design.
Nuclear pulse propulsion, if employed in the 5000-20,000 ton range produces Isp in the 5000-20,000 sec range, at vehicle accelerations in the 2-4 gee range. If used for small ships under 1000 tons, it performance is no better than a solid core NTR. If based on the 1959 fission technology, and on today's more miniaturized fission technology, this is a ready-to-employ technology. EMP and fallout are its drawbacks. Its advantage is lofting truly enormous payloads.
The one that is not ready to employ is the forgotten-to-death gas core NTR technologies (plural). Projections are not reliable precisely because only some benchtop science feasibility experiments were ever done.
There is the "nuclear lightbulb" technology which offers a clean exhaust, and around 1000-1300 sec Isp, at engine T/W perhaps like NERVA or maybe even Timberwind. Who knows? Not ready for anything but more feasibility work.
There is the regeneratively-cooled subcase of the open-cycle gas core NTR that offers 2000-2500 sec Isp at engine T/W maybe as high as 10-to-30, who knows? It has hard radiation in its exhaust, but there is no radioactive core inside the engine after shutdown. Startup would seem to be even slower and more complicated than the solid core NTR, but then again, who knows? This thing could fly as a practical single stage vehicle from the surface. Not ready for anything but more feasibility work.
There is the radiator-cooled subcase of the open-cycle gas core NTR. This one requires a massive hot radiator for the waste heat, because the propellant goes transparent to reacting fireball radiation at the higher power levels. It's a deep space proplsion item only, because the radiator weight drops vehicle accelerations to something in the .01 to .1 gee range. Isp might be in the 6000-10,000 sec range. There is hard radiation in the exhaust. Not ready for anything but more feasibility work.
That last one was one of the 1969-vintage preferred propulsion selections for the then-planned 183 (edit: correct typo "183" to "1983") manned Mars mission at NASA. The other two were NERVA and hydrazine-NTO chemical. They not only cancelled Apollo in 1972, they also cancelled all the nuclear propulsion work. (edit: by the time of Apollo's cancellation, the planned Mars mission had been pushed back to the 1987 opposition.)
Except for Timberwind, nothing has been done since. On any of them.
GW
Last edited by GW Johnson (2016-09-08 09:33:03)
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
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Antius,
It sounds like there's a practical application for the high temperature graphene radiators that NASA is working on:
And the advanced CFRP tanks that ESA is working on (requires autoclaving and liner):
And the advanced CFRP tanks that Boeing is working on (no autoclaving and integrated skirt; not sure about lined / unlined):
Composites World Article on NASA-Boeing Composite Cryotanks
Maybe we can also use the properties of graphene sheet arrangement to make the backing for the radiator an insulator. We're going to need the light weight high temperature radiators and light weight composite tanks to offset the mass of the engine.
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A slight variation on the NTR idea would be a hybrid approach. Basically a two-stage, sea launched mega-rocket scaled to deliver thousands of tonnes of payload to orbit, thereby achieving huge economy of scale and the lift capability needed to deliver colony components.
The lower stage would be nuclear powered; containing what is essentially a boiling water reactor. Water would be stored in a pressurised steel tank and would bleed through the boiling water reactor, exiting through a nozzle. With a chamber exit (pre-nozzle) temperature of 400°C and operating pressure of 200bar, sea level exhaust velocity would be 1379m/s and vacuum exhaust velocity would be 1641m/s. The stage would achieve a total delta-V of ~1000m/s, before releasing the second stage and deploying its drag chute. The reactor would then power down to perhaps 30% of take-off power and use a combination drag and residual thrust to slow sea impact speeds to tolerable levels. After landing, decay heat removal can be accomplished through a natural convection loop dumping heat into sea water through the outer skin of the stage. The rocket system is based upon classic LWR technology and might even use the same fuel elements as commercial BWRs, although fuel enrichment may be different.
The upper stage would consist of a pressure-fed LOX/H2 propulsion system, and would mass about 25,000 tonnes, some 2100 tonnes of which are payload. The stage would be expendable, with propellant tanks and components cannibalised in orbit, providing raw materials for space manufacturing. As the stage will be firing in near vacuum conditions, chamber pressure will be low (perhaps a few bar) and cooling could probably be achieved through use of ablative linings that are sprayed onto the inside of the chamber and nozzle. Liquid hydrogen has such a low heat of evaporation that it can easily self-pressurise. LOX would likely require a heat exchanger loop drawing heat from the engine chamber. Combustion instability is minimised through the use of easily vaporised propellants. Due to the size of the engine, relatively thick ablative linings could be engineered with tolerable weight penalty. Upper stage propellant tanks would be glass-fibre composite and would have internal structure ready for use as a space station module. Engine nozzle would be HSLA steel.
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Ablatives? Meaning you wash it out and recoat for every flight? That's what they had to do with the shuttle SRB's.
Not sure what would work for this hybrid nuke-chemical thing. What worked in the solids was fiber-reinforced EPDM rubber as a case liner, and silica phenolic-and-graphite for the nozzle structure. The graphite is needed for the throat: the pressure differential through that throat eroded silica phenolic char out of there to fast. All that stuff is quite heavy, but essential. (I'm talking real, dense, structural graphite, not some glorified charcoal.)
If the burn is long, EPDM may not be good enough to cover exposed case/chamber walls. There is the silicone material DC-93-104, which forms a much harder surface char layer. More expensive, castable instead of bladdered processing.
I have doubts about a low chamber pressure being adequate. The only thing the nozzle expansion needs is ratio, true (that's CF). But the combustion chamber? No! There are very definite pressure effects. Isp = CF c* / gc and w = P At gc / c* both apply, as well as F = P At CF.
Your Isp depends in part upon your chamber c*, which in turn is a power function of chamber pressure (c* = const * P^ exponent), usually at exponents in the vicinity of 0.1. Going from 1000 psia down to 300 psia is a significant c* loss, on the order of reduction factor 0.89, or an 11% loss. The lower you go, the worse it gets. At 100 psi it is .79 of the 1000 psia value, or a loss of 21%. This is as true of liquids as it is solids.
Only steel strength/weight lowered the chamber pressure of the SRB's to around 900 psia. In the smaller tactical missile motors, we'd typically use 2000-3000 psia as our operating pressure. At 2000 psia, your c* is around factor 1.07 that of 1000 psia, for a 7% gain in chamber performance, not even including the effects of the longer bell available upon nozzle performance (bigger CF).
I think I will include a chapter on booster rocket ballistics in my ramjet book. Seems pertinent enough.
GW
Last edited by GW Johnson (2016-09-08 13:48:34)
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
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Ablatives? Meaning you wash it out and recoat for every flight? That's what they had to do with the shuttle SRB's.
Not sure what would work for this hybrid nuke-chemical thing. What worked in the solids was fiber-reinforced EPDM rubber as a case liner, and silica phenolic-and-graphite for the nozzle structure. The graphite is needed for the throat: the pressure differential through that throat eroded silica phenolic char out of there to fast. All that stuff is quite heavy, but essential. (I'm talking real, dense, structural graphite, not some glorified charcoal.)
If the burn is long, EPDM may not be good enough to cover exposed case/chamber walls. There is the silicone material DC-93-104, which forms a much harder surface char layer. More expensive, castable instead of bladdered processing.
I have doubts about a low chamber pressure being adequate. The only thing the nozzle expansion needs is ratio, true (that's CF). But the combustion chamber? No! There are very definite pressure effects. Isp = CF c* / gc and w = P At gc / c* both apply, as well as F = P At CF.
Your Isp depends in part upon your chamber c*, which in turn is a power function of chamber pressure (c* = const * P^ exponent), usually at exponents in the vicinity of 0.1. Going from 1000 psia down to 300 psia is a significant c* loss, on the order of reduction factor 0.89, or an 11% loss. The lower you go, the worse it gets. At 100 psi it is .79 of the 1000 psia value, or a loss of 21%. This is as true of liquids as it is solids.
Only steel strength/weight lowered the chamber pressure of the SRB's to around 900 psia. In the smaller tactical missile motors, we'd typically use 2000-3000 psia as our operating pressure. At 2000 psia, your c* is around factor 1.07 that of 1000 psia, for a 7% gain in chamber performance, not even including the effects of the longer bell available upon nozzle performance (bigger CF).
I think I will include a chapter on booster rocket ballistics in my ramjet book. Seems pertinent enough.
GW
Many thanks GW. I wasn't aware of the direct significance of chamber pressure, but will upgrade my model. From simple calculations that I have made, I believe that for a pressure-fed upper stage to work with an acceptable mass ratio, tank pressures need to be low, as pressure is directly proportional to weight for any pressure vessel. From what you have described, chamber pressure c* is proportional to the tenth power of pressure, so it would appear that low pressure is way to go, but there would also appear to be an optimum compromise, beyond which reducing pressure further begins to degrade the mass ratio through lower performance and increasing chamber weight. I don't know what point that would be reached, it would be interesting to work it out.
My choice of ablatives was based on the assumption of an expendable upper stage. Hence, there would be no stripping out. This only really works if the stage is cheap to build, I.e. made of steel and really simple, the sort of thing that can be built like a liberty ship. We would be taking it apart and using its components for something else when we reach orbit, I.e. using its tanks as station modules and what remains of the organic liner as a chemical feedstock. So we are reusing it in a different way. Returning it to Earth makes limited sense, given the energy invested in delivering it to orbit.
I don't imagine that ablatives would be an option at high chamber pressures and they would seem to work better in big rather than small chambers. Convective heat transfer to chamber walls is a function of Reynolds number, which in turn is a function of density and chamber diameter. Chamber diameter is huge in this case, but at low chamber pressure, density is lower. Ultimately the abrasion rate is limited by the thermal conductivity of the ablative. For a huge chamber, even a thick lining would not add too much to the weight, as surface area scales to the square of chamber radius, but the weight of the vessel scales to the cube of radius. So it would appear to work better for a big chamber and beneath a certain size abrasion would distort chamber and throat geometry too much anyway.
The lower stage in this case is a nuclear steam rocket. The tanks, chamber and nozzle can be made from carbon steels which remain sufficiently strong at 400C to avoid the need for any cooling. The water is preheated in the tanks, which reduces the heating demand on the reactor and provides a degree of self-pressurisation.
Last edited by Antius (2016-09-08 14:57:35)
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I wonder if pure fusion pulse units could be developed by explosively collapsing a magnetic field around the fusion charge? Surround a lithium deuteride pellet with a shell of superconducting material, and surround the entire assembly with an explosive charge. Connect the superconducting shell to the ship via thin metal cables. Onboard the ship, pulses of power are generated by allowing a uranium hexaflouride gas to go super-prompt critical in a chamber. The plasma escapes through a conducting tube, generating an enormous surge of electric power. The surge electrifies the superconductor through the connecting cables (also converting them to superconducting plasma) milliseconds before the explosive charge is triggered, crushing the superconducting shell onto the fusion charge. As the charge cuts the field lines, eddy currents heat it to plasma and fusion begins.
For this to work, the ship would need to be equipped with a power supply capable of generating huge electrical pulses. Some form of pulsed gaseous fission reactor would appear to be the only thing capable of generating the required power within an acceptable weight budget, by converting fission fragment KE directly into electricity in pulses delivering many gigawatts of power in millisecond durations. To transfer the energy, the pulse units must be connected to ship by wires.
Only a small portion of the fusion charge is likely to consumed in this way. If the pulse units each deliver an explosive power equivalent to 10 times their weight in TNT then the vehicle will greatly outperform any chemical rocket.
Last edited by Antius (2016-09-08 15:55:48)
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I'm sorry, c* is better at higher pressures, not lower. I thought that's what I wrote. This shows up in theoretical thermochemistry, and a similar compounding effect shows up in experimental c* efficiency.
From my ancient Pratt & Whitney Vest Pocket Handbook, c* for LOX-LH2 is quoted as 7950 fps at 1000 psia and r=4.0, and 7840 fps at 100 psia and r=4.5. The r = mass ratio LOX to LH2. The change in r clouds the issue a bit, but I get c*=7950*(p/1000)^0.006 for those data. That's theoretical thermochemical, and when you include c* efficiencies modeled the same way, the exponent is closer to 0.010 or thereabouts.
Maybe this exponent is a little lower with LOX-LH2, maybe higher with the solids. That's why I say somewhere near 0.010 exponent, not 10. If I wrote 0.10, I was wrong. It looks like I did that, though.
As to second stage engine ablative insulation, what you will have after the burn is essentially nothing but a carbon char adhering to the chamber wall. This is due to the thermal cookout wave moving through the hardware, reaching effect in the dozens of seconds after the burn is over. It's pretty easy to wash out with a high-pressure stream of water, but there's not a lot of use for this kind of porous "charcoal".
GW
GW Johnson
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"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
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I'm sorry, c* is better at higher pressures, not lower. I thought that's what I wrote. This shows up in theoretical thermochemistry, and a similar compounding effect shows up in experimental c* efficiency.
From my ancient Pratt & Whitney Vest Pocket Handbook, c* for LOX-LH2 is quoted as 7950 fps at 1000 psia and r=4.0, and 7840 fps at 100 psia and r=4.5. The r = mass ratio LOX to LH2. The change in r clouds the issue a bit, but I get c*=7950*(p/1000)^0.006 for those data. That's theoretical thermochemical, and when you include c* efficiencies modeled the same way, the exponent is closer to 0.010 or thereabouts.
Maybe this exponent is a little lower with LOX-LH2, maybe higher with the solids. That's why I say somewhere near 0.010 exponent, not 10. If I wrote 0.10, I was wrong. It looks like I did that, though.
As to second stage engine ablative insulation, what you will have after the burn is essentially nothing but a carbon char adhering to the chamber wall. This is due to the thermal cookout wave moving through the hardware, reaching effect in the dozens of seconds after the burn is over. It's pretty easy to wash out with a high-pressure stream of water, but there's not a lot of use for this kind of porous "charcoal".
GW
Thanks GW. It is difficult to find clear information on this in the simple text books I have. By 'thermochemical effects', I presume this relates to Chamber Characteristic Length, L*? I.e. at lower pressures, combustion takes longer and characteristic length increases. I can see how that might be a problem in relatively small engines with chamber length <1m in diameter, especially using liquid hydrocarbon fuels which need to vaporise before they can burn. But this would seem to be less of an issue if the chamber dimensions are much larger than typical values for characteristic length.
Convective heat losses to the walls would appear to scale with the Reynolds number, with exponents <1.0. This means that halving chamber pressure will reduce heat transfer rate by slightly less than half. And radiation heat losses from the combustion gases would appear to remain constant with chamber temperature, regardless of pressure.
The second effect looks as if it is more significant than the first and would result in efficiency losses at low chamber pressure even in vacuum. Your exponent of 0.01 would seem to correlate well with heat losses due to convective heat loss effects to the throat and walls, which scale to pressures with an exponent between 0.8 and 0.95, depending which correlations are used. As these are typically only a few percent of total heat generation rate in the engine, a pressure efficiency increase with an exponent of 0.01 sounds about right. The only way to mitigate this is with regenerative cooling.
Clearly high pressure is better from an engine efficiency point of view. But halving combustion chamber pressure almost halves the dry weight of the entire rocket. With the pressure-efficiency exponent being so small, it still looks to me as if low combustion pressure engines provide a better engineering trade-off than high pressure ones. The main detriment so far as I can see would be the weight of the combustion chamber, which will not scale down linearly with reducing pressure due to the need to remove transferred heat. What are your thoughts?
Last edited by Antius (2016-09-09 08:21:26)
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The "thermochemical" I was referring to is two-fold. One is that chamber temperature is a function of chamber pressure, in NASA ODE or any other program. Lower pressures are lower temperatures. That part is just theoretical thermochemistry.
Temperature and predicted properties go into c*. The efficiency of achieving that c* in an actual piece of hardware is also a function of chamber pressure, and decreases at lower pressures. This is entirely empirical, something determined from actual test firings. But for near-stoichiometric mixtures, it's usually pretty high: above maybe 98%.
There is a third effect as well, one that derives from experimental firings. The "optimum" value of r is also a function of chamber pressure. It usually takes a little more oxidizer (bigger r) at the lower pressures to achieve an optimized engine design.
Heat transfer should be significantly affected by chamber pressure through the gas densities (1st power effect on Reynolds number), the slightly-lower temperatures at lower pressures are a minor effect, usually.
For a single-burn design, ablatives are definitely the way to go. That's been true in missile work for over half a century now.
For a steel case that does not get hot, and only gets used once, your hoop stress needs to "cover" your max expected operating pressure by some suitable margins below ultimate, as long as the inelastic strain incurred does not damage seals, insulation, or any other internal structures (such as the case-bonded propellant grain in a solid). For multiple other uses after the burn, you may need bigger margins, and inelastic strain may be inappropriate. If so, you may need to stay a suitable margin below yield during your burn.
That set of design trades sets your wall thickness as a function of pressure. You trade that against performance available vs that same pressure.
GW
Last edited by GW Johnson (2016-09-09 11:22:06)
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
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another topic where we are talking about the old nuclear work that if we had continued it would be more than ready now...
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