Also Yangel was an advocate for storables, and used them in his R-16 missile, which was initially developed as a fall-back to the R-7 but ended up being the Soviets' first practical ICBM. Korolev and his deputy Mishin were the big proponents of kerolox in the Soviet space programme, so it looks like it was more or less OKB-1 against everyone else. By the mid 1960s it is perhaps debatable how much of the opposition to kerolox for launch vehicles was down to engineering reasons and how much was due to personal antagonisms and political rivalries.
Red Star: A Soviet Lunar Landing
- Thread starter SpaceGeek
- Start date
Also Yangel was an advocate for storables, and used them in his R-16 missile, which was initially developed as a fall-back to the R-7 but ended up being the Soviets' first practical ICBM. Korolev and his deputy Mishin were the big proponents of kerolox in the Soviet space programme, so it looks like it was more or less OKB-1 against everyone else. By the mid 1960s it is perhaps debatable how much of the opposition to kerolox for launch vehicles was down to engineering reasons and how much was due to personal antagonisms and political rivalries.
The N1 ITTL was first capable of 75,000 Kg to LEO. Then in the mid-70's, to support the Soviet Lunar Base, it's LEO payload has been upped to 85,000 Kg along with LOX/LH2 Block S & R stages. This is still below what the OTL N1 was supposed to be capable of (95,000 Kg then 105,000 Kg), but it works.
Nixonshead and E of Pi explain this better than I can.
He fall deep as his protector Nikita Khrushchev was removed from power, by Leonid Brezhnev
in OTL he got little support because his Proton rocket worked while N1 explode and N11 remain paper dream.
here the UR-500 never fly means that TKS and ALMAZ not fly also.
the Korolyov Design Bureau had some proposals like Soyuz 7K-VI Zvezda, there answer on USAF MOL
and bigger Soyuz capsule like L3M, 7K-SM, Zarya and ACRV, what give space for maximum, 6 cosmonaut.
So the base N1 has non-cryogenic kerosine to start with, then starts cooling it as part of the early 70s upgrades, 24 engines to start with (staying at 24 engines), the 70s upgrades also include the S & R stages (as they are presented in Encyclopedia Astronautica, I assume). What are the stages in the original and the upgraded version capable of in this TL? I am wondering how much fuel they have, ISP and thrust provided...
With regards to Chelomei, one of the things I was wondering is if his character would be predisposed to tinkering with H2/Lox or solid rocket fuel mixes.
Yes, I know, I am asking alot about Chelomei - I am curious how his character and career would evolve by being defeated by Korolev. As I understand it, Korolev certainly took it hard when Chelomei beat him OTL.
fasquardon
Chelomei fall was partly his own fault, next to be protege of Khrushchev
was he very arrogant with politbureau member, like Ustinov and Marshals responsible for ICBM.
This nasty joke from the time explain allot
next to that were intrigue and power struggle about "Who runs the Soviet Space Program"
a story worthy to be a novel by Leo Tolstoy
as Khrushchev was removed from office the fall of Chelomei starts
the only thing that stop hitting the ground was the Proton rocket that worked while N1 not
Almaz ended up as Salut space station for Soyuz space craft...
As I said at apparently offensive length before
and so will repeat briefly here, the N-1 stages as given in Encyclopedia Astronautica seemed perfectly capable of lifting 75 tons to low orbit, even with the first stage having "only" 24 engines. In particular the very high thrust of even that many engines seemed more than sufficient to lift the whole stack up quite briskly, considering that as given by EA the stages added up to about 2700 tons altogether.
Furthermore, it is hard to tinker with the bottom three stages given their geometrical arrangement. Shrinking one or both of the second and third stages (B and V) would throw their geometrical continuity with the A stage out of whack, due to the general spheres-in-a-cone setup. One might scale the three taken together up or down as a unit but adjusting the relationship between them is tricky!
I have interpreted what Bahamut-255 has said thus far to mean:
1) the bottom 3 stages, A-B-V, are quite close to what Mark Wade cites in Encyclopedia Astronautica. Their dry weight might be somewhat greater, if Wade is taking into account the extreme measures taken OTL to lighten them, and perhaps the mass of propellant is somewhat less, again if Wade factors in the insane chilling stratagem.
2) these 3 stages are just enough to deliver 75 tons to low Earth orbit, presumably at the more or less standard Soviet preferred inclination of 51 degrees (to avoid flying over China, all sources say). I have worked out to my satisfaction that these 75 tons are enough to send the Soyuz-LOK to lunar orbit, on a 50 ton G stage for TLI, leaving the fifth, D stage, enough margin that it can move a Soyuz which itself has plenty of fuel to escape Lunar orbit, into and back out of Lunar orbit and back to Earth.
Now 1) might be more marginal than I thought, because first of all the stages might as noted be heavier dry and contain less fuel. Also there is the Soviet practice of trying to avoid the need for ullage engines by starting the next stage while the one before is running; this means some of the theoretical delta-V of the new stage is wasted.
And there is the question of what happens if an engine goes out, meaning if the KORD control system works right it will shut down the opposite engine as well, so we go from 24 engines down to 22 or perhaps even just 20! Again I believe the design as given above has lots of thrust to spare and the launch can continue.
Note that on this page it is pointed out that the full complement of 30 engines was meant to burn for only 25 seconds in the OTL design--after that 6 engines would be shut down anyway. And 10 seconds before burnout the plan was to shut down all but six of them! I don't know how much good that would have done in terms of avoiding excessive acceleration; I guess that's mainly so that the first stage is providing minimal acceleration to allow the second stage to fire properly.
The outcome of an early shutdown of two or 4 engines, out of 24, would be to draw out the burn--fuel will be gulped down more slowly so the remaining engines burn longer. If it weren't for gravity losses this would make no serious difference since the rocket would still reach the same delta-V, just a bit later--but of course the A stage is mainly fighting gravity and this gives gravity more seconds in which to deduct from the net velocity achieved.
I suppose mission planning makes the pessimistic assumption that a pair of engines will shut down and plans accordingly, which means there will be an unused margin of fuel not burned in the third stage if all goes well.
A lot depends on just when the shutdown happens. I believe though that the N-1 can probably survive a dual engine shutdown right on launch and still continue, as long as another pair doesn't shut down too soon after that. And toward the end of burnout, I'd think as many as eight might be offline and the mission still need not abort. These are just intuitive guesses! But remember the A stage is 2/3 the mass of the whole stack; when it nears empty we'd want to be shutting down engines anyway just to avoid excessive G load.
------
A little fooling around with the Silverbird Calculator suggests to me that if we take the stage masses, ISP, and thrusts Wade gives us, and first deduct both mass and vacuum thrust from the A stage for deleting 6 engines, we can then increase the mass of each empty stage (of the first three that is) by 15 percent, and then lower the mass of fuel by 15 percent, and still hit the 75 tons to orbit target. I doubt they increased the mass of fuel they got into the tanks by chilling it almost to freezing by as much as 15 percent! Bear in mind, most of the fuel mass is liquid oxygen--I don't know how dramatic a density variation you can get in LOX by varying the temperature between near-boiling and near-freezing, but I doubt it is in the ballpark of 15 percent. And maintaining temperatures near freezing for LOX rather than near boiling would be a hell of a challenge. It isn't so difficult to freeze kerosene I guess but again I doubt one gets a 15 percent increase in density. So that leaves all the larger margins for stage weight to increase before a 75 ton payload becomes impossible.
and so will repeat briefly here, the N-1 stages as given in Encyclopedia Astronautica seemed perfectly capable of lifting 75 tons to low orbit, even with the first stage having "only" 24 engines. In particular the very high thrust of even that many engines seemed more than sufficient to lift the whole stack up quite briskly, considering that as given by EA the stages added up to about 2700 tons altogether. Furthermore, it is hard to tinker with the bottom three stages given their geometrical arrangement. Shrinking one or both of the second and third stages (B and V) would throw their geometrical continuity with the A stage out of whack, due to the general spheres-in-a-cone setup. One might scale the three taken together up or down as a unit but adjusting the relationship between them is tricky!
I have interpreted what Bahamut-255 has said thus far to mean:
1) the bottom 3 stages, A-B-V, are quite close to what Mark Wade cites in Encyclopedia Astronautica. Their dry weight might be somewhat greater, if Wade is taking into account the extreme measures taken OTL to lighten them, and perhaps the mass of propellant is somewhat less, again if Wade factors in the insane chilling stratagem.
2) these 3 stages are just enough to deliver 75 tons to low Earth orbit, presumably at the more or less standard Soviet preferred inclination of 51 degrees (to avoid flying over China, all sources say). I have worked out to my satisfaction that these 75 tons are enough to send the Soyuz-LOK to lunar orbit, on a 50 ton G stage for TLI, leaving the fifth, D stage, enough margin that it can move a Soyuz which itself has plenty of fuel to escape Lunar orbit, into and back out of Lunar orbit and back to Earth.
Now 1) might be more marginal than I thought, because first of all the stages might as noted be heavier dry and contain less fuel. Also there is the Soviet practice of trying to avoid the need for ullage engines by starting the next stage while the one before is running; this means some of the theoretical delta-V of the new stage is wasted.
And there is the question of what happens if an engine goes out, meaning if the KORD control system works right it will shut down the opposite engine as well, so we go from 24 engines down to 22 or perhaps even just 20! Again I believe the design as given above has lots of thrust to spare and the launch can continue.
Note that on this page it is pointed out that the full complement of 30 engines was meant to burn for only 25 seconds in the OTL design--after that 6 engines would be shut down anyway.
And 10 seconds before burnout the plan was to shut down all but six of them! I don't know how much good that would have done in terms of avoiding excessive acceleration; I guess that's mainly so that the first stage is providing minimal acceleration to allow the second stage to fire properly.The outcome of an early shutdown of two or 4 engines, out of 24, would be to draw out the burn--fuel will be gulped down more slowly so the remaining engines burn longer. If it weren't for gravity losses this would make no serious difference since the rocket would still reach the same delta-V, just a bit later--but of course the A stage is mainly fighting gravity and this gives gravity more seconds in which to deduct from the net velocity achieved.
I suppose mission planning makes the pessimistic assumption that a pair of engines will shut down and plans accordingly, which means there will be an unused margin of fuel not burned in the third stage if all goes well.
A lot depends on just when the shutdown happens. I believe though that the N-1 can probably survive a dual engine shutdown right on launch and still continue, as long as another pair doesn't shut down too soon after that. And toward the end of burnout, I'd think as many as eight might be offline and the mission still need not abort. These are just intuitive guesses!
But remember the A stage is 2/3 the mass of the whole stack; when it nears empty we'd want to be shutting down engines anyway just to avoid excessive G load.------
A little fooling around with the Silverbird Calculator suggests to me that if we take the stage masses, ISP, and thrusts Wade gives us, and first deduct both mass and vacuum thrust from the A stage for deleting 6 engines, we can then increase the mass of each empty stage (of the first three that is) by 15 percent, and then lower the mass of fuel by 15 percent, and still hit the 75 tons to orbit target. I doubt they increased the mass of fuel they got into the tanks by chilling it almost to freezing by as much as 15 percent! Bear in mind, most of the fuel mass is liquid oxygen--I don't know how dramatic a density variation you can get in LOX by varying the temperature between near-boiling and near-freezing, but I doubt it is in the ballpark of 15 percent. And maintaining temperatures near freezing for LOX rather than near boiling would be a hell of a challenge. It isn't so difficult to freeze kerosene I guess
but again I doubt one gets a 15 percent increase in density. So that leaves all the larger margins for stage weight to increase before a 75 ton payload becomes impossible.
Could it be possible to make it so that the kerosene actually does freeze, and is melted and burned a little at a time?
I mean, you're using LOX anyway; couldn't you also cool the kerosene down to those kinds of temperatures?
Just a little brainstorm here.
I mean, you're using LOX anyway; couldn't you also cool the kerosene down to those kinds of temperatures?
Just a little brainstorm here.
Freezing kerosene would give you all the disadvantages of a solid rocket and all the disadvantages of a liquid rocket plus a few special ones of its own...
Like how would you melt the kerosine at the rate you wanted? Where would you get the thermal energy from?
fasquardon
Like how would you melt the kerosine at the rate you wanted? Where would you get the thermal energy from?
fasquardon
Aaaargh! I suppose you could freeze kerosene; the point is there is little to no point in doing so!
I've spent considerable time trying to find out just how much either LOX or kerosene would contract before freezing, and have yet to come up with a definite answer for either. But I am more convinced than ever--not much.
Kerosene apparently will freeze around -73 C or 200 K. So there is less than 100 degree range between its freezing point and typical "room temperature."
The volumetric coefficient of gasoline at 20 °C is 950 per million degrees K, so that's actually significant--over a 90 degree span that's 8.55 percent. But organic liquids have by far the greater coefficients in general. I have yet had no luck in finding a quote for the volumetric coefficient of LOX but I imagine it is far less than gasoline's, and the thermal range between its freezing and boiling point is also quite low, a matter of 45 degrees or so. Since by volume a ker-LOX engine uses twice as much oxygen as hydrocarbon, it will be the lower percentage variation of LOX that dominates.
8 percent for the hydrocarbon (if kerosene resembles gasoline in this respect) is a lot, but there is no point in loading in 1.08 times the kerosene if you can't also load in 1.08 times the oxygen. Well that's not strictly true; one can always run the engine extra fuel-rich, it will lower the ISP but raise the thrust. But overall the amount of mass increase possible by choosing the ultra-coldest temperature for each liquid can't be nearly as much as 8 percent. When we are dealing with 1800 tons, as in the N-1 A stage, even a few percent add up to quite a mass--5 percent would be 90 tons after all!
That's well and good, but by the same token any loss of control of the temperature will mean the extra propellants come flowing out of the over-tight volumes they've been shoehorned into. That's what I meant by regarding a launch hold as disastrous under the circumstances. (Therefore with this harebrained chilled fuel strategy, the launch controllers are under pressure to avoid holds, cross their fingers, and order the launch on schedule, warnings be damned!)
Now given that, on second thought I suppose the strategy of going ahead and actually freezing the fuels might not be as crazy as I first thought it was. Most substances, unlike water, do not expand on freezing but shrink further--but only a little bit. The coefficient of expansion is much lower for solids than liquids, so we aren't going to be able to squeeze in more material by deep-freezing it. But we would gain some margin for delays before the stuff melts and starts expanding again at a serious rate--not only would it be necessary to heat each kilogram of frozen stuff up to the melting point, but then there would be some heat of fusion to input as well.
As it happens the N-1's spherical tanks lend themselves better to this extreme strategy than other tanks might.
Most liquid fuel rocket engines, at least high-efficiency, high power ones like those used here, involve a coolant loop of either the fuel or oxidant running around the combustion chamber and at least the upper part of the nozzle--otherwise the waste heat from the reaction would melt the materials they are made of. I can see an alternative though--what if we used an inert gas with a very broad thermal range, like say helium, in a closed cycle?
We could run cold (compressed gaseous, at high pressure to achieve high density) helium along the rocket exteriors to cool the material, then route the now very hot and even higher pressure gas first through a turbine--this takes some of the edge of the heat off and usefully reuses some of the waste heat to drive the pumps. Then the still-very-hot gas is diverted into two loops, a small one to melt a suitable amount of kerosene and a bigger one for the oxygen. Each parallel path of hot helium goes up to the tank in question, and loops around the bottom of each sphere, with the tubes at such a spacing and of such a size that the bottom of each tank gets heated by the right amount to melt off just the amount of each propellant we need each second; these flow down to the engines each still about as cold as their melting points plus a bit. There the helium from each tank-melting loop flows back, still pretty hot, to heat exchangers that finally cool the helium back down to a cold temperature to close the cycle and once again cool the engine, while the fuels are vaporized as they blast into the engines.
It's tricky; the question is, is the amount of heat we need to absorb from the engines in line with the amount of heat we need to melt the fuels? I gather that generally speaking the limit of how hot a propellant combination can be burned is when whichever propellant we choose for engine coolant is being completely used up as coolant, so the answer would be, yes, pretty closely in line--consider that here we get to use both propellants, both deeply chilled, as coolants rather than being forced to choose one or the other.
If it turns out we actually need more heat than the engine cooling provides to keep the rate of fuel melting up, we can always have an auxiliary combustion chamber that sacrifices some of the propellants just to generate the additional heat. That will lower the efficiency of the engine since we expend those potential propellants without their serving to generate thrust, or at any rate they generate thrust less efficiently, so the ratio of thrust to mass flow needed to generate it is lowered.
Another advantage of this scheme to consider is that if the propellants are indeed frozen solid, then we might be better able to control their sloshing. When the tanks are completely full the stuff can swirl about all it likes and it won't upset the balance of the rocket (though dynamic forces on the tank walls might break them) but when they are partially emptied, sloshing can throw the balance of the rocket off. But if they are solid chunks, with say a light network of cables running through them to the tank wall, their solidity will hold them in place! The tricky thing here is that perhaps instead of smoothly melting off, chunks will stay stubbornly attached to the cables and fail to melt off and fall to the bottom where the heat can melt them. Maybe we can electrically heat the cables as the level of fuel drops to their height?
By going indirect like this with a third fluid we lose some efficiency, but if helium is the third fluid we know we can avoid many thermochemical headaches that using oxygen instead directly in the chamber cooling and in the drive turbine would inflict on us.
So these are advantages of what at first looked like an utterly wrongheaded idea to me, hence my cry of pain!
All this said, if the name of the game is to lower the overall dry weight of the rocket relative to the fuel it contains--I suspect we still have a loser here. I've proposed helium plumbing, containing hot helium at temperatures comparable to those that melt engines and at high pressures demanding considerable strength to contain it (strength compromised by the high temperature too!) running all the way up to the upper kerosene tank, and looping around the bottoms of both tanks, then running back to the engines to be run through heat exchangers there so we can cool it back to the low temperature needed to cool the engines. All of this will add weight; it seems doubtful to me that the few percent we gain in propellant we can contain in a given volume will offset the mass of the helium plumbing.
So if we refrain from actually freezing the fluids and go back to just chilling them, we have no thermal margin.
For the kerosene anyway, one of the major advantages of using it rather than more powerful hydrogen is avoiding the headaches of cryogenics. (Another, probably really the bigger advantage, is density--kerosene would be something like 10-12 times denser than the equivalent mass of LH2, so the tankage is smaller--and if we avoid shenanigans involved with chilling it, we don't need to insulate the tanks either) Oxygen on the other hand must be cryogenic anyway, and it will boil off unless we go beyond merely liquefying it to freeze or near-freeze it--and even that just buys us time, it will heat up and come to a boil anyway after that. I suppose that it is managed on the launching pad by venting off the boil-off and pumping in more LOX, or by accepting that some of the tankage must go to contain some only briefly, before it does boil off.
Perhaps then a hybrid strategy, where we do freeze the LOX, and use something like my helium loop with helium-driven turbines and heat exchangers on the oxygen only? The oxygen is nearly 2/3 the propellant mass and would be doing the lion's share of any cooling, directly or indirectly, anyway, and with the N-1 design the tank is right there at the bottom of the stage next to the engines, so the plumbing to connect the helium coming from each engine to the bottom of the oxy tank would be short. Meanwhile we let the kerosene be at ambient temperature and simply pump it into the engines.
Since as far as I can tell the oxygen would not be much reduced in volume by freezing it versus keeping it near boiling temperatures anyway, this means none of the heroic foolishness of trying to squeeze in more fuel that the OTL N-1 operations were committed to; the fuel is what it is, at ambient temperatures, and the designers just have to live with it.
---
Given the math I did last night, I figure then that if Bahamut's N-1 can only lift just 75 tons to orbit, then the main difference between the OTL figures Wade gives and this timelines lies in a somewhat heavier structure for each stage. I would guess, given what I've managed to learn about the thermal volumetric coefficients of the propellants and the temperature ranges available, that for all their efforts at chilling the fuels they couldn't manage more than a 5 percent increase over the more reasonable temperatures I presume the authors' design plans around. So there is margin for the empty stages to be as much as 20 percent heavier than OTL, which would seem reasonable when we consider that corners were cut to an extreme extent there OTL.
I might run the figures in Silverbird and present them for consideration here tonight.
I've spent considerable time trying to find out just how much either LOX or kerosene would contract before freezing, and have yet to come up with a definite answer for either.
But I am more convinced than ever--not much. Kerosene apparently will freeze around -73 C or 200 K. So there is less than 100 degree range between its freezing point and typical "room temperature."
The volumetric coefficient of gasoline at 20 °C is 950 per million degrees K, so that's actually significant--over a 90 degree span that's 8.55 percent. But organic liquids have by far the greater coefficients in general. I have yet had no luck in finding a quote for the volumetric coefficient of LOX but I imagine it is far less than gasoline's, and the thermal range between its freezing and boiling point is also quite low, a matter of 45 degrees or so. Since by volume a ker-LOX engine uses twice as much oxygen as hydrocarbon, it will be the lower percentage variation of LOX that dominates.
8 percent for the hydrocarbon (if kerosene resembles gasoline in this respect) is a lot, but there is no point in loading in 1.08 times the kerosene if you can't also load in 1.08 times the oxygen. Well that's not strictly true; one can always run the engine extra fuel-rich, it will lower the ISP but raise the thrust. But overall the amount of mass increase possible by choosing the ultra-coldest temperature for each liquid can't be nearly as much as 8 percent. When we are dealing with 1800 tons, as in the N-1 A stage, even a few percent add up to quite a mass--5 percent would be 90 tons after all!
That's well and good, but by the same token any loss of control of the temperature will mean the extra propellants come flowing out of the over-tight volumes they've been shoehorned into. That's what I meant by regarding a launch hold as disastrous under the circumstances. (Therefore with this harebrained chilled fuel strategy, the launch controllers are under pressure to avoid holds, cross their fingers, and order the launch on schedule, warnings be damned!
)Now given that, on second thought I suppose the strategy of going ahead and actually freezing the fuels might not be as crazy as I first thought it was.
Most substances, unlike water, do not expand on freezing but shrink further--but only a little bit. The coefficient of expansion is much lower for solids than liquids, so we aren't going to be able to squeeze in more material by deep-freezing it. But we would gain some margin for delays before the stuff melts and starts expanding again at a serious rate--not only would it be necessary to heat each kilogram of frozen stuff up to the melting point, but then there would be some heat of fusion to input as well.As it happens the N-1's spherical tanks lend themselves better to this extreme strategy than other tanks might.
Most liquid fuel rocket engines, at least high-efficiency, high power ones like those used here, involve a coolant loop of either the fuel or oxidant running around the combustion chamber and at least the upper part of the nozzle--otherwise the waste heat from the reaction would melt the materials they are made of. I can see an alternative though--what if we used an inert gas with a very broad thermal range, like say helium, in a closed cycle?
We could run cold (compressed gaseous, at high pressure to achieve high density) helium along the rocket exteriors to cool the material, then route the now very hot and even higher pressure gas first through a turbine--this takes some of the edge of the heat off and usefully reuses some of the waste heat to drive the pumps. Then the still-very-hot gas is diverted into two loops, a small one to melt a suitable amount of kerosene and a bigger one for the oxygen. Each parallel path of hot helium goes up to the tank in question, and loops around the bottom of each sphere, with the tubes at such a spacing and of such a size that the bottom of each tank gets heated by the right amount to melt off just the amount of each propellant we need each second; these flow down to the engines each still about as cold as their melting points plus a bit. There the helium from each tank-melting loop flows back, still pretty hot, to heat exchangers that finally cool the helium back down to a cold temperature to close the cycle and once again cool the engine, while the fuels are vaporized as they blast into the engines.
It's tricky; the question is, is the amount of heat we need to absorb from the engines in line with the amount of heat we need to melt the fuels? I gather that generally speaking the limit of how hot a propellant combination can be burned is when whichever propellant we choose for engine coolant is being completely used up as coolant, so the answer would be, yes, pretty closely in line--consider that here we get to use both propellants, both deeply chilled, as coolants rather than being forced to choose one or the other.
If it turns out we actually need more heat than the engine cooling provides to keep the rate of fuel melting up, we can always have an auxiliary combustion chamber that sacrifices some of the propellants just to generate the additional heat. That will lower the efficiency of the engine since we expend those potential propellants without their serving to generate thrust, or at any rate they generate thrust less efficiently, so the ratio of thrust to mass flow needed to generate it is lowered.
Another advantage of this scheme to consider is that if the propellants are indeed frozen solid, then we might be better able to control their sloshing. When the tanks are completely full the stuff can swirl about all it likes and it won't upset the balance of the rocket (though dynamic forces on the tank walls might break them) but when they are partially emptied, sloshing can throw the balance of the rocket off. But if they are solid chunks, with say a light network of cables running through them to the tank wall, their solidity will hold them in place! The tricky thing here is that perhaps instead of smoothly melting off, chunks will stay stubbornly attached to the cables and fail to melt off and fall to the bottom where the heat can melt them. Maybe we can electrically heat the cables as the level of fuel drops to their height?
By going indirect like this with a third fluid we lose some efficiency, but if helium is the third fluid we know we can avoid many thermochemical headaches that using oxygen instead directly in the chamber cooling and in the drive turbine would inflict on us.
So these are advantages of what at first looked like an utterly wrongheaded idea to me, hence my cry of pain!
All this said, if the name of the game is to lower the overall dry weight of the rocket relative to the fuel it contains--I suspect we still have a loser here. I've proposed helium plumbing, containing hot helium at temperatures comparable to those that melt engines and at high pressures demanding considerable strength to contain it (strength compromised by the high temperature too!) running all the way up to the upper kerosene tank, and looping around the bottoms of both tanks, then running back to the engines to be run through heat exchangers there so we can cool it back to the low temperature needed to cool the engines. All of this will add weight; it seems doubtful to me that the few percent we gain in propellant we can contain in a given volume will offset the mass of the helium plumbing.
So if we refrain from actually freezing the fluids and go back to just chilling them, we have no thermal margin.
For the kerosene anyway, one of the major advantages of using it rather than more powerful hydrogen is avoiding the headaches of cryogenics. (Another, probably really the bigger advantage, is density--kerosene would be something like 10-12 times denser than the equivalent mass of LH2, so the tankage is smaller--and if we avoid shenanigans involved with chilling it, we don't need to insulate the tanks either) Oxygen on the other hand must be cryogenic anyway, and it will boil off unless we go beyond merely liquefying it to freeze or near-freeze it--and even that just buys us time, it will heat up and come to a boil anyway after that. I suppose that it is managed on the launching pad by venting off the boil-off and pumping in more LOX, or by accepting that some of the tankage must go to contain some only briefly, before it does boil off.
Perhaps then a hybrid strategy, where we do freeze the LOX, and use something like my helium loop with helium-driven turbines and heat exchangers on the oxygen only? The oxygen is nearly 2/3 the propellant mass and would be doing the lion's share of any cooling, directly or indirectly, anyway, and with the N-1 design the tank is right there at the bottom of the stage next to the engines, so the plumbing to connect the helium coming from each engine to the bottom of the oxy tank would be short. Meanwhile we let the kerosene be at ambient temperature and simply pump it into the engines.
Since as far as I can tell the oxygen would not be much reduced in volume by freezing it versus keeping it near boiling temperatures anyway, this means none of the heroic foolishness of trying to squeeze in more fuel that the OTL N-1 operations were committed to; the fuel is what it is, at ambient temperatures, and the designers just have to live with it.
---
Given the math I did last night, I figure then that if Bahamut's N-1 can only lift just 75 tons to orbit, then the main difference between the OTL figures Wade gives and this timelines lies in a somewhat heavier structure for each stage. I would guess, given what I've managed to learn about the thermal volumetric coefficients of the propellants and the temperature ranges available, that for all their efforts at chilling the fuels they couldn't manage more than a 5 percent increase over the more reasonable temperatures I presume the authors' design plans around. So there is margin for the empty stages to be as much as 20 percent heavier than OTL, which would seem reasonable when we consider that corners were cut to an extreme extent there OTL.
I might run the figures in Silverbird and present them for consideration here tonight.
So, I did it. I took Wade's figures, deducted the 6 engines from the A stage and their weight, took the new vacuum thrust that implied for the A stage (Silverbird wants vacuum figures, I suppose it discounts thrust and ISP to account for atmospheric pressure for you) and then lowered the propellant masses Wade's figures imply by 5 percent, and raised the dry structure masses he gives by 20 percent. And, bearing in mind that Bahamut cautioned us the ascent profile for the N-1 required that the shroud covering the Soyuz and its launch escape rocket would remain with the payload all the way to orbit (dumb as that seems!) I input a shroud mass of 4500 kg and that it would separate at 625 seconds, which is the sum of the burn times Wade gives for the three stages ABV.
Here's the stats for these three stages, as thus modified:
Block A
No. engines: 24
Gross mass: 1,809,500 kg
Unfuelled mass: 147,000 kg
Propellants 1,662,500 kg
Height: 30.10 m (98.70 ft).
Diameter: 10.30 m (33.70 ft).
Span: 16.90 m (55.40 ft).
Thrust: 41008.60 kN. vacuum
Specific impulse: 318 s.
Specific impulse sea level: 284 s.
Burn time: 125 s. (156.25)
Block B
No Engines: 8.
Gross mass: 546,590 kg
Unfuelled mass: 66,840 kg
Propellants 479,750 kg
Height: 20.50 m (67.20 ft).
Diameter: 6.80 m (22.30 ft).
Span: 9.80 m (32.10 ft).
Thrust: 14,039.98 kN (3,156,313 lbf).
Specific impulse: 346 s.
Burn time: 120 s.
Block V
No. engines: 4
Gross mass: 182,690 kg
Unfuelled mass: 16,440 kg
Propellants 166,250 kg
Height: 14.10 m (46.20 ft).
Diameter: 4.80 m (15.70 ft).
Span: 6.40 m (20.90 ft).
Thrust: 1,608.00 kN (361,492 lbf).
Specific impulse: 353 s.
Burn time: 370 s.
Adding up to
2,538,859,500 kg
with 75 ton payload to orbit and 4500 kg shroud/LES
Here's the Silverbird output:
If we jettison the shroud at 140 seconds after launch, that is 15 seconds into the second stage burn the calculator raises the total mass to
Looking at the OTL 90 ton to orbit, single launch plan for both LOK and LK in one vehicle, I can see better why they'd plan to keep the shroud all the way to orbit (and it would be a bigger shroud too!)--the shroud covers both the Soyuz and the LK, the latter in particular is an especially delicate craft designed for vacuum and Lunar operations. OTL Apollo kept the LM under a conical shroud until TLI was complete, at which point the CSM would undock from it--the shroud for the LM being also a load-bearing structure that carried the CSM. Only after TLI was accomplished would the two moon ships be in zero G and free to separate from the TLI stack. Presumably the OTL N-1 stack had to provide similar support for the Soyuz and protection for the LK, so perhaps it would even have been needed to be included in the TLI boost!
Here on the other hand we have no such situation; the Soyuz is presumably mounted directly on the De stage, and the shroud merely provides aerodynamic protection, as well as being the attachment of the launch escape rocket that pulls the two upper Soyuz modules loose in case of launch failure. I therefore took the heavier SAS (Systema Avarinovo Spassenya, the Russian term for the launch escape rocket) shroud combination of 4.5 tons from this page. I suppose, given that it needs to escape from a much bigger rocket, the SAS part of the mass needs to be greater and the shroud too strengthened. And unfortunately I have no idea just how far into launch the Semyorka-launched Soyuzes jettisoned these--oh wait, look here--apparently they ditched the escape rocket at 113 seconds after launch, before the booster stages burned out, then the shroud separately at 158, still before the core of the first stage burned out. 140 seconds that I chose was based on the idea that you jettison all this stuff after the first stage is burned out, a little bit into the second stage burn, as per Apollo.
(By then you see, not only has the vast majority of the dangerously potentially explosive fuel been burned off leaving only a relatively small fraction--also, the craft as a whole is high in the atmosphere in very thin air, so an explosion would not carry the same punch--and it is moving through that thin air at high supersonic speeds, so the blast tends to be left behind. At that point, figured Apollo mission planners, the Apollo Service Module main engine could manage the job of separating the CM from the catastrophe briskly enough. So too with the N-1, although the second stage and the 3 above it still pack quite a wallop, the same considerations apply--though the Soyuz engine is perhaps not as powerful, it also has a much lesser mass to shift, so I suppose at roughly the same point in launch the SAS and the shroud are disposable. Maybe the Soyuz needs the shroud a bit longer because unlike Apollo it is completely unstreamlined, so we need to wait a bit to get into even thinner air before discarding it--but at any rate we don't need the escape rocket, and clearly the standard Soyuz shroud was designed to peel off on its own without needing the SAS to pull it.)
Clearly then if there was a need OTL to retain the shroud all the way to orbit, it related to the situation of the double LOK-LK stack and not to Soyuz operations as such. Since ITTL the manned launch contains a Soyuz only, clearly we can jettison the shroud long before reaching orbit, though perhaps a bit later than I guessed, and the dang thing is probably considerably heavier than 4500 kg!
But even if the latter factor outweighs the former and cuts our orbital payload down, we can clearly achieve the 75 ton goal with this rocket. Perhaps the Silverbird calculator fails to account for aspects of Soviet engineering or the possibility of engine failures and so overestimates the payload--but even so, small reductions in the empty masses of the first three stages should close the gap.
Your N-1 is a good rocket! Don't discount it!
Here's the stats for these three stages, as thus modified:
Block A
No. engines: 24
Gross mass: 1,809,500 kg
Unfuelled mass: 147,000 kg
Propellants 1,662,500 kg
Height: 30.10 m (98.70 ft).
Diameter: 10.30 m (33.70 ft).
Span: 16.90 m (55.40 ft).
Thrust: 41008.60 kN. vacuum
Specific impulse: 318 s.
Specific impulse sea level: 284 s.
Burn time: 125 s. (156.25)
Block B
No Engines: 8.
Gross mass: 546,590 kg
Unfuelled mass: 66,840 kg
Propellants 479,750 kg
Height: 20.50 m (67.20 ft).
Diameter: 6.80 m (22.30 ft).
Span: 9.80 m (32.10 ft).
Thrust: 14,039.98 kN (3,156,313 lbf).
Specific impulse: 346 s.
Burn time: 120 s.
Block V
No. engines: 4
Gross mass: 182,690 kg
Unfuelled mass: 16,440 kg
Propellants 166,250 kg
Height: 14.10 m (46.20 ft).
Diameter: 4.80 m (15.70 ft).
Span: 6.40 m (20.90 ft).
Thrust: 1,608.00 kN (361,492 lbf).
Specific impulse: 353 s.
Burn time: 370 s.
Adding up to
2,538,859,500 kg
with 75 ton payload to orbit and 4500 kg shroud/LES
Here's the Silverbird output:
If we jettison the shroud at 140 seconds after launch, that is 15 seconds into the second stage burn the calculator raises the total mass to
Looking at the OTL 90 ton to orbit, single launch plan for both LOK and LK in one vehicle, I can see better why they'd plan to keep the shroud all the way to orbit (and it would be a bigger shroud too!)--the shroud covers both the Soyuz and the LK, the latter in particular is an especially delicate craft designed for vacuum and Lunar operations. OTL Apollo kept the LM under a conical shroud until TLI was complete, at which point the CSM would undock from it--the shroud for the LM being also a load-bearing structure that carried the CSM. Only after TLI was accomplished would the two moon ships be in zero G and free to separate from the TLI stack. Presumably the OTL N-1 stack had to provide similar support for the Soyuz and protection for the LK, so perhaps it would even have been needed to be included in the TLI boost!
Here on the other hand we have no such situation; the Soyuz is presumably mounted directly on the De stage, and the shroud merely provides aerodynamic protection, as well as being the attachment of the launch escape rocket that pulls the two upper Soyuz modules loose in case of launch failure. I therefore took the heavier SAS (Systema Avarinovo Spassenya, the Russian term for the launch escape rocket) shroud combination of 4.5 tons from this page. I suppose, given that it needs to escape from a much bigger rocket, the SAS part of the mass needs to be greater and the shroud too strengthened. And unfortunately I have no idea just how far into launch the Semyorka-launched Soyuzes jettisoned these--oh wait, look here--apparently they ditched the escape rocket at 113 seconds after launch, before the booster stages burned out, then the shroud separately at 158, still before the core of the first stage burned out. 140 seconds that I chose was based on the idea that you jettison all this stuff after the first stage is burned out, a little bit into the second stage burn, as per Apollo.
(By then you see, not only has the vast majority of the dangerously potentially explosive fuel been burned off leaving only a relatively small fraction--also, the craft as a whole is high in the atmosphere in very thin air, so an explosion would not carry the same punch--and it is moving through that thin air at high supersonic speeds, so the blast tends to be left behind. At that point, figured Apollo mission planners, the Apollo Service Module main engine could manage the job of separating the CM from the catastrophe briskly enough. So too with the N-1, although the second stage and the 3 above it still pack quite a wallop, the same considerations apply--though the Soyuz engine is perhaps not as powerful, it also has a much lesser mass to shift, so I suppose at roughly the same point in launch the SAS and the shroud are disposable. Maybe the Soyuz needs the shroud a bit longer because unlike Apollo it is completely unstreamlined, so we need to wait a bit to get into even thinner air before discarding it--but at any rate we don't need the escape rocket, and clearly the standard Soyuz shroud was designed to peel off on its own without needing the SAS to pull it.)
Clearly then if there was a need OTL to retain the shroud all the way to orbit, it related to the situation of the double LOK-LK stack and not to Soyuz operations as such. Since ITTL the manned launch contains a Soyuz only, clearly we can jettison the shroud long before reaching orbit, though perhaps a bit later than I guessed, and the dang thing is probably considerably heavier than 4500 kg!
But even if the latter factor outweighs the former and cuts our orbital payload down, we can clearly achieve the 75 ton goal with this rocket. Perhaps the Silverbird calculator fails to account for aspects of Soviet engineering or the possibility of engine failures and so overestimates the payload--but even so, small reductions in the empty masses of the first three stages should close the gap.
Your N-1 is a good rocket! Don't discount it!
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Shevek, One thing I think you miss in the helium heating system is the need of something to stir the tanks to make sure that the O2 and kerosine melt evenly. Also, consider what might happen if a chunk of frozen whatever gets into a line and blocks it?
It is a very cool idea for a system, but I am pretty sure it is one that would tend towards catastrophic failure...
fasquardon
It is a very cool idea for a system, but I am pretty sure it is one that would tend towards catastrophic failure...
fasquardon
That could fairly easily be prevented by including some sort of strainer assembly over the input to the fuel\oxidiser lines.
Oh, good point. And they already needed such filters in the N1 too.
fasquardon
The Soviets had been enjoying three years alone with the Moon, without the NASA J-Class Apollo missions to challenge their status as leading space fairing nation. Despite this, Mishin and the engineers at TsKBEM knew that this era was coming to an end, and thing were about to change radically. The Americans were about to launch their month-long Apollo 20 orbital mission while in just a year's time NASA's LESA lunar habitat would begin the process of building a permanent lunar base. Meanwhile the technically inferior LK Shelter lumbered through development, slowed down by multiple delays.
L3-9 would be the last of the "traditional" Soviet lunar missions. All following missions would employ the LK shelter for long duration missions and deploy larger crews. This was Mishin's Apollo 19, the end and beginning of an era. The end of short, lunar sorties, our species first preliminary reconnaissance of the Moon and beginning of our first outpost. Unlike prior missions this was focused much more on investigation of the site that had been chosen (though not publicly released) for the first Soviet lunar base.
Vladislav Volkov was proud to be the commander of this mission. Having flown previously multiple times in LEO he could finally have his chance of reaching for the Moon. Alongside him was Flight Engineer Georgy Dobrovolsky. Unlike Volkov this was Dobrovolsky's first spaceflight of any kind, he would go straight from clinging to the Earth to soaring hundreds of thousands of km to lunar orbit without any prior spaceflight experience. He was excited to say the least.
The Beginning of the transit was normal enough, but in keeping with this rather unwanted tradition of Soviet spaceflight, the toilet malfunctioned again. This was a problem encountered since the early days of Vostok and despite all the great marvels of Soviet spaceflight had still not been eliminated completely. While the crew of the Zarya enjoyed a more conventional toilet similar to what might be encountered on Earth, the Soyuz crews were condemned to use a small, imperfect system that made the environment, to quote Bykovsky on his Vostok 5 flight "unpleasant".
Space Poop aside, the Lunar orbit insertion went normally having been performed countless times before by prior crews and before the Russian duo knew it, they were coming up on their landing site. As had been performed successfully by both Apollo and L3 missions alike, Volkov made a surface rendezvous with Clauvius Crater. As the tiny LK made it's final descent the seismic readings picked up the crashing Block D before Salyut's (meaning "Union" in Russian) landing gear touched down with a clunk. Volkov stepping out onto the surface could already see the landing gear and flag left behind by a previous mission. While no man had visited the site before, that didn't mean it hadn't seen visitors. Irina Solovyova had set foot just several hundred feet away years prior, bringing new meaning to the phrase "where no man had gone before". The Red flag itself was perfectly fine although knocked over by the dust kicked up from L3-6's lunar ascent. Dusting it off and raising it up was a proud moment. The beginnings of a lunar base had started.
Unlike previous missions (intended to garner public attention and prestige milestones) this mission was focused primarily on continuing the scientific work done by Solovyova, mainly evaluating it as a potential location for a future permanent base. A Lunakhod rover and Luna lander had reinforced the view left behind by L3-6, that this was the ideal location for a lunar base. Arthur C. Clarke would later be asked about this turn of events in a suspicious light.
After two intensive EVAs, bagging rocks and performing multiple environmental surveys aimed at learning significantly more about the area Volkov left, 24 hours and 3 seconds after he arrived. The handshake in orbit was an enthralling time for the both of them. While 24 hours doesn't seem like long for most it does after performing such an intense exploration in such a short time period. Dobrovolsky wasn't just twiddling his thumbs while coasting, silently, alone, in orbit. Instead he was busily mapping the lunar surface with a suit of various cameras and instruments, all in high resolution. After a day of pure relaxation as they enjoyed the view of the Sun, Earth and Moon from space the cosmonauts initiated the all-important "TEI" burn. Another three days of coasting brought humanity's first program of Manned lunar exploration to a close. But while the "conventional" L3 program finished it was really the beginning of a new and exciting lifetime for the Soyuz and LK as they embarked on longer expeditions in space and time. But only if they could beat the Americans...
L3-9 would be the last of the "traditional" Soviet lunar missions. All following missions would employ the LK shelter for long duration missions and deploy larger crews. This was Mishin's Apollo 19, the end and beginning of an era. The end of short, lunar sorties, our species first preliminary reconnaissance of the Moon and beginning of our first outpost. Unlike prior missions this was focused much more on investigation of the site that had been chosen (though not publicly released) for the first Soviet lunar base.
Vladislav Volkov was proud to be the commander of this mission. Having flown previously multiple times in LEO he could finally have his chance of reaching for the Moon. Alongside him was Flight Engineer Georgy Dobrovolsky. Unlike Volkov this was Dobrovolsky's first spaceflight of any kind, he would go straight from clinging to the Earth to soaring hundreds of thousands of km to lunar orbit without any prior spaceflight experience. He was excited to say the least.
The Beginning of the transit was normal enough, but in keeping with this rather unwanted tradition of Soviet spaceflight, the toilet malfunctioned again. This was a problem encountered since the early days of Vostok and despite all the great marvels of Soviet spaceflight had still not been eliminated completely. While the crew of the Zarya enjoyed a more conventional toilet similar to what might be encountered on Earth, the Soyuz crews were condemned to use a small, imperfect system that made the environment, to quote Bykovsky on his Vostok 5 flight "unpleasant".
Space Poop aside, the Lunar orbit insertion went normally having been performed countless times before by prior crews and before the Russian duo knew it, they were coming up on their landing site. As had been performed successfully by both Apollo and L3 missions alike, Volkov made a surface rendezvous with Clauvius Crater. As the tiny LK made it's final descent the seismic readings picked up the crashing Block D before Salyut's (meaning "Union" in Russian) landing gear touched down with a clunk. Volkov stepping out onto the surface could already see the landing gear and flag left behind by a previous mission. While no man had visited the site before, that didn't mean it hadn't seen visitors. Irina Solovyova had set foot just several hundred feet away years prior, bringing new meaning to the phrase "where no man had gone before". The Red flag itself was perfectly fine although knocked over by the dust kicked up from L3-6's lunar ascent. Dusting it off and raising it up was a proud moment. The beginnings of a lunar base had started.
Unlike previous missions (intended to garner public attention and prestige milestones) this mission was focused primarily on continuing the scientific work done by Solovyova, mainly evaluating it as a potential location for a future permanent base. A Lunakhod rover and Luna lander had reinforced the view left behind by L3-6, that this was the ideal location for a lunar base. Arthur C. Clarke would later be asked about this turn of events in a suspicious light.
After two intensive EVAs, bagging rocks and performing multiple environmental surveys aimed at learning significantly more about the area Volkov left, 24 hours and 3 seconds after he arrived. The handshake in orbit was an enthralling time for the both of them. While 24 hours doesn't seem like long for most it does after performing such an intense exploration in such a short time period. Dobrovolsky wasn't just twiddling his thumbs while coasting, silently, alone, in orbit. Instead he was busily mapping the lunar surface with a suit of various cameras and instruments, all in high resolution. After a day of pure relaxation as they enjoyed the view of the Sun, Earth and Moon from space the cosmonauts initiated the all-important "TEI" burn. Another three days of coasting brought humanity's first program of Manned lunar exploration to a close. But while the "conventional" L3 program finished it was really the beginning of a new and exciting lifetime for the Soyuz and LK as they embarked on longer expeditions in space and time. But only if they could beat the Americans...
Nice buildup to the next installment!
Something I wonder is if the extra fuel in the descent stages from prior landings will still be useable years later?
fasquardon
Something I wonder is if the extra fuel in the descent stages from prior landings will still be useable years later?
fasquardon
The other issue, of course, being that the LK uses the same stage to return to orbit as it does to land--there was no separate "descent stage" tanks and engines that would be left on the surface, just the landing gear (see image here). There could still be residuals left in the Apollo descent stage tanks, but it'd depend on how heat is transited within the stage structure during the lunar day.L3-6 landed in Clauvius in early 1973
L3-9 landed in Clauvius in mid 1975
By then any fuel left behind by the L3-6 descent stage would be gone.
Love the brainstorming though
SpaceGeek and Bahamut, do you have any thoughts on Shevek's calculations on the N1's capabilities? Interested in what you both think!
Me, I don't understand the machinery nearly well enough to make any estimations.
fasquardon
Me, I don't understand the machinery nearly well enough to make any estimations.
fasquardon