ApolloHoax.net
Apollo Discussions => The Reality of Apollo => Topic started by: Bob B. on March 15, 2012, 06:24:03 PM
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I wonder what the Prof would make of the images of the Black Arrow launch.... (see attached)
And that was on earth as well.
The Black Arrow was fueled by kerosene and hydrogen peroxide, wasn't it?
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The Black Arrow was fueled by kerosene and hydrogen peroxide, wasn't it?
Yes, H202/RP-1
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I wonder why the plume is transparent despite the use of kerosene fuel? Maybe the mixture ratio didn't have to be rich as in most rockets.
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I wonder why the plume is transparent despite the use of kerosene fuel? Maybe the mixture ratio didn't have to be rich as in most rockets.
Hydrogen peroxide decomposition produces two molecules of water for each molecule of O2, so the sooty highly incandescent particles would be a bit more diluted in the exhaust. Might also use a more oxidizer-rich mixture due to the fact that just decomposing peroxide contributes some to the thrust, while unburned fuel contributes nothing. Lower temperatures might also be a factor in allowing oxidizer-rich mixtures. Hard information would be nice...
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I later read that the invisible exhaust is attributed to the large amount of water from the decomposition of H2O2. I can understand that but it was also my understanding that most rockets are run fuel-rich to avoid an oxidizing environment that could erode the combustion chamber and nozzle, and with RP-1 fuel that would leave some unreacted carbon to glow when it hits atmospheric O2.
I would think that the atomic oxygen released by the decomposition of H2O2 would be at least as corrosive as any oxidizer, but maybe not.
I've noticed in the video from Deltas, Falcons and other RP-1 burning rockets that as they rise in the atmosphere, their plumes broaden out and become less bright, but you can still see what look like random black streaks leaving the nozzle. I assume they're made of unburned carbon that no longer has a chance to burn and glow in atmospheric O2, but I'm not sure. If you look very closely at some of the Apollo 16mm landing films you will also see those little black streaks on occasion, which I would also attribute to unreacted carbon from the carbon-containing Aerozine-50 fuel.
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I would think that the atomic oxygen released by the decomposition of H2O2 would be at least as corrosive as any oxidizer, but maybe not.
HTP's been used as a monopropellant in numerous rockets (not often...performance is poor and it's expensive, hazardous, prone to decomposition in the presence of any contamination, etc). Any atomic oxygen has probably recombined to molecular hydrogen by the time it mixes with the kerosene and burns, so you're basically using a mixture of steam and oxygen as your oxidizer, which is probably a good bit cooler burning and less of a corrosion issue than injecting pure liquid oxygen.
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I seem to recall reading somewhere that the Black Arrow used 85% hydrogen peroxide, though I don't know the mixture ratio of oxidizer to fuel. That's going to produce a great deal of water in the exhaust. The temperature is also much lower than when LOX is used. A LOX/RP-1 engine might have a combustion chamber temperature of around 3300o C, while a HTP/RP-1 (HTP=high test peroxide) engine might operate at about 2400o C. I'm by no means an expert of the incandescence of carbon, but I have to believe the lower temperature would decrease the effect.
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... it was also my understanding that most rockets are run fuel-rich to avoid an oxidizing environment that could erode the combustion chamber and nozzle
Reducing corrosion might be part of the reason, but I've always understood the reason for operating fuel-rich is to lower the molecular weight of the exhaust gas. Oxidizing the fuel creates heavy molecules -- CO2 is heavier than CO, and H2O is heavier than H2 or H -- which results in a lower exhaust gas velocity. The downside of running fuel-rich is that the temperature is less, which also reduces exhaust velocity. There's a happy medium where we supply adequate oxygen to burn enough of the fuel to get a high temperature, but not too much oxygen that we drive the molecular weight too high. The optimum mixture that achieves the highest exhaust velocity is on the fuel-rich side of a stoichiometric ratio.
By the way, the Russians operate many of their engines oxidizer-rich in the preburner and turbines. I haven't studied their designs enough to know how they've overcome the corrosion issue, but they've apparently figured it out. I think their N2O4/UDMH staged-combustion engines operate oxidizer-rich, though the main combustion chamber is still burning a fuel-rich mixture. They route the oxidizer along with a small amount of fuel to the preburner, and then the oxygen rich gas is routed to the turbines. The turbine exhaust then goes to the combustion chamber where it is combined with the remaining fuel and burned.
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What's the problem with a lower exhaust velocity?
What's the problem with high mass lowering exhaust velocity? That would mean you could get the same thrust for a reduced power.
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so you're basically using a mixture of steam and oxygen as your oxidizer, which is probably a good bit cooler burning and less of a corrosion issue
Steam can be notoriously corrosive, as any power plant operator can tell you.
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so you're basically using a mixture of steam and oxygen as your oxidizer, which is probably a good bit cooler burning and less of a corrosion issue
Steam can be notoriously corrosive, as any power plant operator can tell you.
Water in general as any well operator can tell you.
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What's the problem with a lower exhaust velocity?
What's the problem with high mass lowering exhaust velocity? That would mean you could get the same thrust for a reduced power.
Yes, but only at the cost of greater propellant mass, so it's not usually worth it.
In a chemical rocket, the energy is stored in the same material that becomes the reaction mass, so you might as well store as much energy as possible in that mass so you can expel it as quickly as possible. That's why you always want chemical propellants with the highest specific impulse available, subject only to practical factors like toxicity, density (which affects tank weight), cost, reliability, safety, etc.
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What's the problem with high mass lowering exhaust velocity? That would mean you could get the same thrust for a reduced power.
I just realized you probably actually meant to ask what's the problem with a high molecular weight exhaust lowering exhaust velocity.
Every rocket is a heat engine. However the energy is stored, whether it's in the chemical energy of the propellants themselves or added externally from a nuclear reactor or a electrically-powered arc, it first becomes heat. Then that heat is turned into the mechanical (kinetic) energy of the propellant, and that's what makes it a heat engine.
No heat engine can be 100% efficient in turning heat into mechanical energy for a variety of reasons. In a rocket, one of the factors affecting that conversion efficiency is the molecular weight of the exhaust, the lower the better. This can be such an important factor that sometimes it pays to convert less of the stored chemical energy into heat in the first place, if that permits a more efficient conversion of the heat you do generate into mechanical energy.
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Steam can be notoriously corrosive, as any power plant operator can tell you.
Not as corrosive as high concentrations of oxygen at high temperatures (http://en.wikipedia.org/wiki/Thermal_lance). The Soviets had to work out some metallurgical tricks to handle such conditions, IIRC, while dealing with steam was well understood by that time.
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Preburners are run lean to keep the exhaust gas temperatures low until they reach the thrust chamber, to prevent excess thermal stress on the preburner and turbine components. Those are notoriously difficult to cool.
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Right, I know that keeping a mixture either extremely lean or extremely rich will keep preburner temperatures low by limiting the amount of heat produced and diluting that heat with a large amount of unreacted oxidizer or fuel, respectively. What I don't understand is how a lean mixture could have any kind of advantage over a rich mixture, given that hot oxidizers are even more corrosive than cold oxidizers. Hot hydrocarbons are generally far more benign on metals with the possible exception of embrittlement by free hydrogen.
Maybe the Russians couldn't bring themselves to do something the same way as us western capitalist imperialist stooges.
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Preburners are run lean to keep the exhaust gas temperatures low until they reach the thrust chamber, to prevent excess thermal stress on the preburner and turbine components. Those are notoriously difficult to cool.
Or they are run rich, as is the case with the SSME. In the SSME, most of the hydrogen is run through the preburner (about 70% of the total flow as I recall) with only a small amount of oxygen. In either case, lean or rich, the preburners operate far from the stoichiometric ratio to keep the temperature down. Same thing for a gas generator in an open cycle engine, though I think they almost always run rich.
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Or they are run rich, as is the case with the SSME.
Wow, ignore anything I write before about 9:00 am Utah time. Yes, all American engines I know run their auxiliary or prestage cycles fuel-rich, oxidizer-lean (which is how my pre-caffeinated mind was thinking) in order to keep combustion temperatures low enough for radiative cooling. One big consequence of that in RP-1 fueled engines is coking of the turbine. It's less a problem in hydrogen-fueled engines.
The RP-1 prestage exhaust products are very sooty. It is those sooty carbon species that ignite and/or incandesce in the F-1 flow beginning a few feet below the exit plane. Keep in mind that the Rocketdyne F-1 used the gas-generator exhaust (fuel-rich) as film cooling for the nozzle extension.
Metallurgy hasn't historically been Russian engineering's strong point. I'll have to ask some colleagues who are more familiar with Russian engines why they can run fuel-lean.
I think the SSME ran its preburner absurdly fuel-rich in order to capitalize on some obscure adiabatic expansion properties, but that may be apocryphal.
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What I don't understand is how a lean mixture could have any kind of advantage over a rich mixture, given that hot oxidizers are even more corrosive than cold oxidizers. Hot hydrocarbons are generally far more benign on metals with the possible exception of embrittlement by free hydrogen.
Maybe the Russians couldn't bring themselves to do something the same way as us western capitalist imperialist stooges.
I can't respond with certainty, but one possible explanation is that they're using the fuel as the coolant. I know UDMH is adequate for regenerative cooling but I've never heard of using N2O4 for that purpose. If the fuel is the only adequate source of coolant, then they obviously have to route that through the thrust chamber cooling channels. That may leave them with oxidizer as the only option to power the turbines. Though I don't suppose it's a real show stopper if the fuel was used as both the coolant and the propellant for the turbines, running it in series from the cooling jacket to the preburner.
Another possibility it that there isn't a large enough mass of fuel to power the turbines. Gas generator cycle engines require only a few percent of the fuel to be diverted to the gas generator, but staged combustion engines require a large propellant flow through the preburner and turbines. The large mass flow is required for a couple reasons: (1) higher pump horsepower due to the greater discharge pressure, and (2) less pressure drop across the turbines compared to gas generators. The turbine pressure ratio of a gas generator engine is large because the exhaust is simply dumped overboard. In staged combustion the exhaust from the turbines must remain at a very high pressure because it is then forced into the combustion chamber, thus the turbine pressure ratio is low. In a N2O4/UDMH engine there is more than twice as much oxidizer as there is fuel. Perhaps the turbines require a greater flow rate than the fuel alone can provide.
I honestly don't know the answer. I just thinking out loud and suggesting possibilities.
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I'll have to ask some colleagues who are more familiar with Russian engines why they can run fuel-lean.
If you can do that I'd be very interested in hearing their explanations. Thanks.
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I think the SSME ran its preburner absurdly fuel-rich in order to capitalize on some obscure adiabatic expansion properties, but that may be apocryphal.
Aside from the temperature issue, I can't confirm if there were any other reasons, but you're correct about the mixture being very fuel-rich. I just found an article stating that 76% of the hydrogen flow and 11% of the oxygen flow was injected into the two preburners. The fuel preburner had a mixer ratio of 0.86 and the oxidizer preburner had a mixture ratio of 0.60. This compares to a mixture ratio of 6.03 at the main injector.
Source: http://large.stanford.edu/courses/2011/ph240/nguyen1/docs/SSME_PRESENTATION.pdf
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With LOX/RP1, you have greater volumes of LOX to move than RP1. Running the preburner fuel-rich would further increase the burden on the LOX pumps relative to the RP1 pumps, running them lean might be part of why the Russian engines are so lightweight for their thrust. With LOX/LH2, the situation is reversed.
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I took a look at the book History of Liquid Propellant Rocket Engines by George P. Sutton to see if it said anything about fuel-rich versus oxidizer-rich. The book addressed the issue but didn't provide the answers I was hoping it would:
In the United States and several other countries, the GGs and PBs are usually operated at a fuel-rich mixture ratio. There are only a few exceptions, such as Goddard's GGs, which were oxidizer rich. In the Soviet Union and more recently Russia, the vast majority of GGs and PBs are usually operated at an oxidizer-rich mixture ratio. The exceptions are LOX/LH2 LPREs and a few others, where the gases are fuel rich. There are some reasons for the choice of fuel-rich or oxidizer-rich gases, when driving turbines, but the rationales are usually not very strong, and a new engine could be operated with either one of these gas mixtures.
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With LOX/LH2, the situation is reversed
LH2 is indeed extremely bulky (density 0.07 that of water) but it also has extremely low viscosity that makes it easy to pump.
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Question on 1960's Soviet and USA soft landing of probes on the moon.
Could someone explain to me in plain english how this was accomplished?
The time needed to transmit and receive signals to and from a probe landing on the moon is significant.
To program a device to land via a set program could not take into account any discrepancies for significant alterations in parameters, such as altitude, terrain so on.
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I can't say for some of them, but the first successful soft lander, Luna 9, had a contact probe extending down from the main bus carrying the instrument unit and rocket. Upon contact, the main bus ejected the airbag clad instrument unit, which unfolded like a flower, revealing the camera and antenna. A rather elegant system I must say.
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Question on 1960's Soviet and USA soft landing of probes on the moon.
Could someone explain to me in plain english how this was accomplished?
The time needed to transmit and receive signals to and from a probe landing on the moon is significant.
To program a device to land via a set program could not take into account any discrepancies for significant alterations in parameters, such as altitude, terrain so on.
Could someone explain to me in plain english how this was accomplished?
A seemingly reasonable search for knowledge.
The time needed to transmit and receive signals to and from a probe landing on the moon is significant.
But this statement indicates you have already made some judgment in the matter. Significance is relative to the problem and how much time is needed. The round trip time was about about 3 seconds, a short time relative to a Mars landing for instance. The Russians started landing on Mars in the early 70's.
To program a device to land via a set program could not take into account any discrepancies for significant alterations in parameters, such as altitude, terrain so on.
Why not? This is a claim of fact that you cannot support. Why do you just assume his to be true?
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...the main bus ejected the airbag clad instrument unit, which unfolded like a flower, revealing the camera and antenna. A rather elegant system I must say.
And a system still in use today for Mars landers.
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Could someone explain to me in plain english how this was accomplished?
Well the Rangers did not soft-land. They plummeted into the surface.
The Surveyors soft-landed. The majority of the trip to the Moon was in a translunar coast. At a certain point determined by the ground, based on telemetry, tracking, and a predetermined plan, the spacecraft was turned around and a powerful retrorocket was fired. This rocket slowed the approach from cislunar speeds to an approach speed, timed precisely to leave the spacecraft at a certain altitude above the surface, descending at a certain rate.
Then the terminal descent program kicked in, which merely maintained a constant descent rate until contact.
The time needed to transmit and receive signals to and from a probe landing on the moon is significant.
Correct, which is why no real-time control was attempted from Earth during phases were close timing was critical. The retrofire ignition was time-critical, but it was a singular event. Once initiated, no further ground control was needed. You just send the ignition command 1.3 seconds before you need ignition to occur. After that, the spacecraft guides itself (i.e., maintains the proper attitude).
To program a device to land via a set program could not take into account any discrepancies for significant alterations in parameters, such as altitude, terrain so on.
Correct. The model for retrofire ignition was based on an idealized shape of the Moon, which indeed would not allow for local variations in terrain such as hills or hummocks. Hence the terminal descent program is meant to start relatively high (to account for any local maxima) and descent at a constant rate of descent until it contacts the surface. This rate of descent is fast enough to be sustainable for several seconds, yet slow enough to be within the limits of the landing gear to absorb.
Imagine groping in your dark bedroom for the lightswitch. If you get up from your bed and start walking, you know approximately how far it is to the wall. You may walk briskly for a few steps to cover the distance, then slow down and walk slowly with your arm extended until you touch the wall. You walk slowly enough that it won't hurt when you run into it.
Now you may ask how they would avoid boulders, craters, and other obstacles that may possibly upset the spacecraft. The answer is that they didn't. That was part of the risk of the mission. They accepted the possibility that one leg may land on a large boulder and that the whole thing would tip over. This is why manned landings had a generally higher expectation of success; the pilot could see and avoid obstacles because he is actually there seeing them and piloting the vehicle.
In short, Surveyor's descent was open-loop guidance, Apollo's descent was closed-loop guidance.
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Can anyone briefly summarize the relative advantages and disadvantages of the various turbine-driven rocket engine fuel cycles, i.e., gas generators vs staged combustion vs expander cycle? I've read the Wikipedia articles for each but I don't fully understand why they each have the advantages they do.
In particular, why is a gas generator cycle engine inherently less efficient? The turbine exhaust isn't actually thrown away, it is still ejected downward where it can add a little thrust. In single-engine stages it can even be gimbaled to provide roll control, which I would think is a very useful feature.
I know that the fuel-rich mixture used in most gas generators does represent unburned and therefore wasted fuel, but this can come in handy for nozzle protection as in the F-1, possibly allowing the bulk of the exhaust to be at a hotter temperature than the engine might otherwise tolerate.
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The time needed to transmit and receive signals to and from a probe landing on the moon is significant.
Yes it is, which is why remote control landing was not attempted.
To program a device to land via a set program could not take into account any discrepancies for significant alterations in parameters, such as altitude, terrain so on.
This is not strictly true. Yes it is true that you cannot simply program the vehicle as if it were landing on a perfect sphere. What you can do is a) program it such that it will land safely on a variety of landscapes, and b) design some feedback systems in so that it can in effect 'see' the ground it is trying to land on.
You can achieve a) by programming a rapid initial deceleration then maintaining a speed that will not destroy the probe on impact with the ground no matter if it occurs earlier or later than the ideal altitude (say if it runs into a hummock or a crater).
b) is achieved by the use of systems such a radar units or contact probes, which tell the probe exactly how far away the surface is at any given time and command it to respond accordingly. The Apollo lunar module, for example, used both methods. A radar told the crew how high they were and how far they were descending, and the contact probes on the landing gear told the crew they had touched down. In Apollo this input was output as information on displays and the crew responded to it. It would not require a complicated processor to take those inputs and use them to trigger commands to tell the craft to do something, for example slow down if it is going too fast, or shut down its engine when it is just about to land.
And no, there was no way to avoid unexpected boulders, or the problems of inadvertently landing the craft in the edge of a sheer drop and having it tumble in. That's just the risk of the mission. A few years ago I attended a talk by John Zarencki about the Cassini-Huygens mission, in the run up to the landing of Huygens on Titan. In the Q&A session afterwards I asked about what would happen if the lander hit a rock and flipped over, and his response was the 'I'm glad you asked that' reply of someone who had hoped no-one would find the flaw in his plan. There was indeed, as he told us, no provision for such an eventuality and it was simply considered to be a risk that was acceptable. The mission would still yield usueful information during the descent.
To extend Jay's dark room analogy, you walk with your hand outstretched so that your fingertips hit the wall first. That way the impact is absorbed by your arm and not your face, allowing you to slow down before you hit the wall and do yourself some damage. That's the same principle as a contact probe under the feet of a lander. What that can't help you with, however, is the toybox on the floor that was unexpectedly left there for you to trip over, but to be scanning everything all around you for any obstacles all the way is a huge effort for the desired outcome, so you simply accept that as a risk.
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Forgive my imagination...
To program a device to land via a set program could not take into account any discrepancies for significant alterations in parameters, such as altitude, terrain so on.
In a NASA office, far away and long ago..
Chair: Now, onto the landing procedures. As the lander descends towards the cratered & rocky surface, what will..
ProfM: Um, excuse me.. What did you say?
Chair: You mean about descending, or the cratered, undulating and rock-strewn surface?
ProfM: Craters, you say??? Undulating surface, you say? And there are rocks?
Chair: Well, yes, of course there are - we know that because..
ProfM: No, no no - all of my calculations and procedures were based on the fact that the Moon is a perfectly smooth sphere! After all, I have looked at it through my binoculars - it is clearly smooth.
Chair: I'm sorry Munkin, but that isn't the case..
ProfM: Well, we CANNOT land there, then!! How could we POSSIBLY land on something where we might ... tip over, or land at an angle? I'm sorry, it is impossible. The mission is off.
Me, I think problem-solving should be taught from kindergarten..
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I can't quite figure out from the question - is our Professor merely curious as to *how* they did it, or is he suggesting that it was impossible, so the unmanned missions were faked?
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Can anyone briefly summarize the relative advantages and disadvantages of the various turbine-driven rocket engine fuel cycles, i.e., gas generators vs staged combustion vs expander cycle? I've read the Wikipedia articles for each but I don't fully understand why they each have the advantages they do.
In particular, why is a gas generator cycle engine inherently less efficient? The turbine exhaust isn't actually thrown away, it is still ejected downward where it can add a little thrust. In single-engine stages it can even be gimbaled to provide roll control, which I would think is a very useful feature.
I know that the fuel-rich mixture used in most gas generators does represent unburned and therefore wasted fuel, but this can come in handy for nozzle protection as in the F-1, possibly allowing the bulk of the exhaust to be at a hotter temperature than the engine might otherwise tolerate.
The turbine exhaust of a gas generator doesn't really add all that much thrust; most of the energy is used to drive the turbines. The gas exiting the turbines has relatively low temperature and pressure, so there's not a lot more that can be gotten out of it. Usually the amount of propellant that is routed to the gas generator is about 2% to 7% of the total, so all other things being equal, a gas generator engine should be just a few percent less efficient than staged combustion.
The big difference that I've observed is that staged combustion engines can operate at much higher pressures, yielding considerably higher specific impulse compared to gas generators. The highest chamber pressure I can recall for a gas generator engine is about 96 atmospheres in the RS-68, while some staged combustion engines operate at chamber pressures exceeding 250 atmospheres. This, I think, is really the big advantage of staged combustion.
Running the pumps of a gas generator engine at a higher pressure means more propellant must be routed to the gas generator/turbines, and since this propellant produces little in the way of thrust, increasing the flow harms the efficiency of the engine. We therefore have two competing factors at work -- increasing the chamber pressure increases efficiency, but the resulting greater propellant flow to the gas generator decreases efficiency. These factors have to be balanced, and it seems that the upper limit on the chamber pressure of a gas generator engine is about 100 atmospheres.
In a staged combustion engine, the turbine exhaust, which is still at very high pressure, is routed into the combustion chamber and burned with the remaining propellant to produce thrust. In this design a very large proportion of the propellant can be routed through the preburner and turbines without a significant decrease in efficiency. This allows the engines to operate at very high pressures compared to gas generators. In some designs, as much as 100% of one propellant (fuel or oxidizer) is routed through the preburner/turbines before being injected into the combustion chamber.
The expander cycle has the advantage of staged combustion in that all the propellant produces thrust, but it has the disadvantage of a gas generator in that the operating pressure seems to be limited. The expander cycle works only with cryogenic propellants, most commonly LOX/LH2 though probably also LOX/CH4.
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b) is achieved by the use of systems such a radar units or contact probes, which tell the probe exactly how far away the surface is at any given time and command it to respond accordingly.
And it should be noted that Surveyor was equipped with doppler and altimeter radars to measure descent velocity and altitude.
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It's a variation of the argument we had with fattydash on the old boards regarding the rendezvous with the orbiting CSM, in which he wanted to argue that it all needed to be calculated in advance. In fact it's a combination of a calculated course getting you roughly where you need to be, then an active system updating your information for refining the final approach.
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I can't quite figure out from the question - is our Professor merely curious as to *how* they did it, or is he suggesting that it was impossible, so the unmanned missions were faked?
He's an HB, what do you think? The number of times I can recall an HB being honestly just curious I can probably count on one hand. They're almost always probing for some piece of information they can use to claim hoax, regardless of how innocent they make their inquiries sound.
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Thanks guys for both the question and the answer about gas generators vs staged combustion. I learned something new and interesting.
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Can anyone briefly summarize...
Jay doesn't do "briefly." ;D
In particular, why is a gas generator cycle engine inherently less efficient?
Briefly, thermodynamics. You throw away heat and enthalpy when you dump turbine exhaust overboard, even if the exhaust can be put to productive use in other ways. Feeding it back into the main thermodynamic design captures it more effectively in terms of measured thrust.
The turbine exhaust isn't actually thrown away, it is still ejected downward where it can add a little thrust.
Not typically enough to matter. You gain more by directing it back into explicit thrust-generating processes rather than trying to use its thrust directly. Ironically the turbines in the open-cycle solution can be a little more efficient because their exhaust vents to the ambient. You can extract more mechanical power through conscientious mechanical design of the turbine. The exhaust in a closed-cycle design has to maintain enough residual pressure to feed the next stage where you have to contend against chamber pressure, hence the turbine design has to be a little less efficient.
Consider the difference between marine turbine engines and aircraft turbine engines. The powerplant on a Perry-class frigate in the U.S. Navy is composed of two GE engines of exactly the same type as are used on the Boeing 747 (4, in this case) to provide thrust. It's quite valid to say that a Navy frigate and a passenger airliner use the same engine. However, in the marine case the aft-end turbine is designed extract as much mechanical power as possible, leaving very little power in the flow of the exhaust gas, which is simply vented out the stack. But in the aerospace case, the turbine is design to extract only as much mechanical power from the exhaust stream as is required to operate the engine's compressor and bypass fan and PTO needs of the airframe, and leave as much power in the exhaust stream as possible for producing Newtonian thrust.
In single-engine stages it can even be gimbaled to provide roll control, which I would think is a very useful feature.
Yes, this is often done.
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...the main bus ejected the airbag clad instrument unit, which unfolded like a flower, revealing the camera and antenna. A rather elegant system I must say.
And a system still in use today for Mars landers.
Yes, Oppoertuinity and Spirit, as well as Mars Pathfinder, also used what can be most succinctly described as "lithobreaking" for the final decent.
Ranger probes also used it in a rather terminal fashion.
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Bob & Jay, thanks guys. That helps.
I was curious to know just how much power the turbopumps represent in the big picture, i.e., how much mechanical power the pumps apply to the propellants versus the overall heat rate of the rocket engine.
Turns out it's pretty tiny, for the F-1 at least. Using the available figures for the RP-1 and LOX flow rates, their densities, the combustion chamber pressure and the S-IC tank ullage pressures, I get 10.8 MW for the LOX pump and 6.7 MW for the RP-1 pump, per engine. Using the heating value of RP-1, that's only 0.05% of the total heat rate of the engine of 34.2 GW. And that's for a single F-1 engine! I'm sure the real turbines developed considerably more as I neglected the kinetic energy in the moving propellants and the inefficiency of the pump impellers.
Do those numbers sound about right?
So yeah, using even 2% of your propellant in a gas generator to develop only 0.05% of its energy content does seem rather inefficient, even for a heat engine.
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It occurs to me that if turbines in conventional gas-generator cycle engines must run with inefficiently rich mixtures of the main propellant supply to keep from burning up, why not use some auxiliary fuel that could operate the gas generator more efficiently?
The fuel that comes to mind is hydrazine, or perhaps hydrazine hydrate to lower the temperature if necessary. Catalytically decomposed hydrazine contains no oxygen, just ammonia, nitrogen and hydrogen, so it should be fairly benign on turbine blades. Another advantage is that since hydrazine spontaneously decomposes when it hits a catalyst, the complex bootstrapping processes now necessary to start many large rocket engines could be simplified.
The Shuttle's APUs were powered this way, although they were much too small (135 hp) to drive the turbopumps on a large rocket engine.
Edit to add: I'm not surprised to see this is not an original idea. The Germans thought of it first by using H2O2 and a catalyst to drive the alcohol and LOX turbopumps in the V2.
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Ranger probes also used it in a rather terminal fashion.
Yes, it's always seemed a shame that the only way to soft-land on the moon is to burn even more propellant than you need to get back off the surface again. If only you could land on some sort of net stretched across a big crater. Problem is, it would have to be a rather deep big crater.
The moon's escape velocity is 2.38 km/sec, so that's the absolute minimum speed at which you'd hit the surface on a direct unbraked approach from infinity. If you could tolerate a maximum of 10g's, that would be a deceleration of about 98 m/sec2, so it would take you 24+ seconds to come to a stop.
Okay, that's not so bad with a really well designed couch but during those 24 seconds you'd travel 28.9 km. That's a pretty deep crater.
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It occurs to me that if turbines in conventional gas-generator cycle engines must run with inefficiently rich mixtures of the main propellant supply to keep from burning up, why not use some auxiliary fuel that could operate the gas generator more efficiently?
I haven't studied this issue real close, but I think the disadvantage of using a third propellant, such as H2O2 or N2H4, is a higher inert mass. There are additional tanks, piping, etc. that aren't necessary when we tap off the main propellants. Of course, the main propellant tanks can be a little smaller, but I doubt this offsets the mass of a separate system. I'd also like to see how much power can be obtained from a kilogram of monopropellant versus a kilogram of fuel-rich bipropellant to see if there is an advantage to one versus the other. I should be able to perform these calculations, but I'll have to get back to you.
Edit to add: I'm not surprised to see this is not an original idea. The Germans thought of it first by using H2O2 and a catalyst to drive the alcohol and LOX turbopumps in the V2.
One of the main reasons the Germans used this method is because they had prior experience with it, having used H2O2 decomposition to power torpedos. The Germans also had the ability to manufacture H2O2 in high concentrations (I think about 70% at the time of WWII). The American Redstone also used the method, being that it was a derivative of the V-2. Likewise, many of the early Russian designs used H2O2 to power its turbopumps, including the R-7. I'm pretty sure the present day Soyuz launch vehicle still uses it.
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Ranger probes also used it in a rather terminal fashion.
Yes, it's always seemed a shame that the only way to soft-land on the moon is to burn even more propellant than you need to get back off the surface again. If only you could land on some sort of net stretched across a big crater. Problem is, it would have to be a rather deep big crater.
The moon's escape velocity is 2.38 km/sec, so that's the absolute minimum speed at which you'd hit the surface on a direct unbraked approach from infinity. If you could tolerate a maximum of 10g's, that would be a deceleration of about 98 m/sec2, so it would take you 24+ seconds to come to a stop.
Okay, that's not so bad with a really well designed couch but during those 24 seconds you'd travel 28.9 km. That's a pretty deep crater.
That gives me Bugs Bunny visions.
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I knew I should have made a left turn at Albuquerque Station. :P
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Of course, the main propellant tanks can be a little smaller
Or the main engines can burn a little longer.
I'd also like to see how much power can be obtained from a kilogram of monopropellant versus a kilogram of fuel-rich bipropellant to see if there is an advantage to one versus the other.
Another idea: if the problem is to avoid burning up the turbine blades, why not burn a stoichiometric ratio of the main propellants along with something to cool the flame and produce even more gas -- like maybe water?
The efficiencies I got for gas generators seemed so low that almost anything ought to be an improvement, I'd think.
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Another idea: if the problem is to avoid burning up the turbine blades, why not burn a stoichiometric ratio of the main propellants along with something to cool the flame and produce even more gas -- like maybe water?
The efficiencies I got for gas generators seemed so low that almost anything ought to be an improvement, I'd think.
Reducing the temperature is exactly the opposite of what you want to do to increase efficiency. The maximum efficiency of a heat engine is 1 - Tcold/Thot, and you're talking about reducing Thot to something closer to Tcold. You're also adding a separate tank, plumbing, and pumps for water or some other fluid that you have to carry but which doesn't contribute any energy...if you must reduce operating temperature, you may as well just use some extra oxidizer or fuel.
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Reducing the temperature is exactly the opposite of what you want to do to increase efficiency.
Oh, I know that. The problem is that the turbine blades simply can't withstand the actual flame temperature, so you have no choice but to cool the hot gases before they reach the blades even though that reduces the thermodynamic efficiency of the heat engine.
Whether you mix water with the hot gases or combine the propellants in a highly non-stoichiometric ratio so they don't burn as hot in the first place, either way you are decreasing efficiency. But I don't know enough about turbine design to know which approach reduces the efficiency less.
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I don't know if I'm going about this the right way or not, but I've attempted to compare the effectiveness of several propellants for driving turbines. I've assumed in all cases that the pressure at the turbine inlet is 1000 PSI and and the pressure at the outlet is 50 PSI, for a pressure ratio of 20:1. For each propellant, I calculated the enthalpy of the gases at the inlet and then at the outlet assuming isentropic expansion. Taking the difference in enthalpy, I think, gives the amount of energy available to drive the turbine. If the enthalpy difference is low, that means a larger mass of propellant is needed to produce a given turbine power.
In the first case I took 100% hydrogen peroxide, which when decomposed at 1000 PSI gives a temperature of 1279 K and an enthalpy of -5.5145 MJ/kg. Expanding the gas until the pressure drops to 50 PSI reduces the temperature to 669 K and the enthalpy to -6.5573 MJ/kg. The enthalpy difference is 1.0428 MJ/kg. (Note that concentrations below 100% are less effective propellants.)
Next I did hydrazine. The inlet and outlet temperatures are 911 K and 530 K, and the enthalpy difference is 1.5005 MJ/kg.
Next on my list is fuel-rich LOX and kerosene (for which I used C12H26). I used a mixture ratio of 1.13 because this gives just the right amount of oxygen to oxidize all the carbon to CO, but not enough to oxidize the hydrogen. This mixture yields inlet and outlet temperatures of 1599 K and 1090 K, and the enthalpy difference is 2.0502 MJ/kg. (In practice, other mixture ratios might be common, which I'd have to do further research to determine.)
Finally, I used a stoichiometric mixture of LOX and kerosene with water added. I added enough water to lower the inlet temperature to approximately the same as that of the fuel-rich LOX/kerosene mixture. I ended up with about 59% water by mass. The inlet and outlet temperatures are 1595 K and 925 K, and the enthalpy difference is 1.4233 MJ/kg.
Therefore, if my logic and math are correct, fuel-rich LOX/kerosene is best, requiring the least amount of propellant to drive the turbines. Hydrogen peroxide is the worst, requiring double the mass of LOX/kerosene. Hydrazine and watered-down LOX/kerosene are about equal and lie in the middle.
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Very interesting!
Now I have to go back to the books to make sure I know what "Enthalpy" is and how and why it changes... :-)
Oh, when you did hydrazine, what did you decompose it to? Ammonia and nitrogen, or nitrogen and hydrogen?
And what about the energy necessary to vaporize the LOX? The other propellants are all stored at room temperature.
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Very interesting!
Now I have to go back to the books to make sure I know what "Enthalpy" is and how and why it changes... :-)
I've done some more calculations since yesterday. I wanted to bring the LOX/kerosene combustion temperature down to something comparable to the hydrogen peroxide, so I lowered the mixture ratio from 1.13 to 0.5. This changed the inlet and outlet temperatures to 1270 K and 958 K. It also lowered the enthalpy change to 1.4902 MJ/kg.
Similarly, I increased the amount of water in the watered-down scenario to about 64%. This changed the inlet and outlet temperatures to 1275 K and 707 K, and lowered the enthalpy change to 1.1460 MJ/kg.
With these revised numbers we see that LOX/kerosene doesn't look as good as it did before, but it's still better than H2O2 and about equal to hydrazine. Of course if we lower the mixture ratio further, it gets closer and closer to H2O2.
Oh, when you did hydrazine, what did you decompose it to? Ammonia and nitrogen, or nitrogen and hydrogen?
I included H, H2, N2 and NH3 in the products and let my computer program (STANJAN) sort out the equilibrium mixture. On the inlet side I had almost complete dissociation of the ammonia (about 98%). On the outlet side, because of the lower temperature, some of the ammonia reformed to where I had about 87% dissociation.
It's my understanding that it's possible to vary the design of the catalyst chamber to control the amount of ammonia dissociation. About 30%-40% is the minimum that can be maintained, which results in a hotter temperature of about 1400 K. This is desirable in thrusters because less dissociation produces a higher specific impulse. However, I've read that a higher percentage dissociation is typically allowed in gas generators to keep the temperature lower.
Unfortunately, with the computer program I'm using, I have to control over how much ammonia dissociation there is. I think that if I could reduce the dissociation and get a higher temperature, we'd find that the change in enthalpy would likely be considerably better than LOX/kerosene.
And what about the energy necessary to vaporize the LOX? The other propellants are all stored at room temperature.
That's been taken into account. I reduced the heat in the gas by the amount needed to vaporize the LOX and bring it up to the same starting temperature as the other propellants.
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With these revised numbers we see that LOX/kerosene doesn't look as good as it did before, but it's still better than H2O2 and about equal to hydrazine. Of course if we lower the mixture ratio further, it gets closer and closer to H2O2.
It also (in the case of LOX/RP1 rockets, anyway) uses the stuff you already have large, mass-efficient tanks full of and which you already need to pump around.
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Very interesting!
Now I have to go back to the books to make sure I know what "Enthalpy" is and how and why it changes... :-)
Oh, when you did hydrazine, what did you decompose it to? Ammonia and nitrogen, or nitrogen and hydrogen?
And what about the energy necessary to vaporize the LOX? The other propellants are all stored at room temperature.
Enthalpy is the heat of reaction. Or maybe it's better termed the change in chemical energy of the system since an exothermic reaction has negative energy.
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Watching this thread develop is fascinating.
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Watching this thread develop is fascinating.
It is interesting to talk about this stuff, but we're so far off topic now that we should have started a new thread.
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Introduction to enthalpy as it relates to the fluid-cycle design of rocket engines.
http://exploration.grc.nasa.gov/education/rocket/enthalpy.html
This is the e-book for Sutton and Biblarz, the standard work. The thermodynamics chapter is skimmable without having to buy the book.
http://books.google.com/books?id=pFktw0GYSX8C
Yes, this is an awesome thread, and I'm sure LunarOrbit wouldn't object to splitting it off into the Reality forum.
Bob, what program are you using? Is it something I can rewrite to give you control over the ammonia dissociation?
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Introduction to enthalpy as it relates to the fluid-cycle design of rocket engines.
http://exploration.grc.nasa.gov/education/rocket/enthalpy.html
Please allow me to add the following about isentropic compression/expansion.
http://www.grc.nasa.gov/WWW/K-12/airplane/compexp.html
If it will help, I'll provide a quick example of the calculations I've been doing.
Let's do the calculation for atomic hydrogen, H. As mentioned earlier, I assumed a turbine pressure ratio of 20:1, and let's assume our turbine inlet temperature is 1200 K. Atomic hydrogen has a specific heat ratio, γ, of 1.667 and it's heat capacity (http://webbook.nist.gov/cgi/cbook.cgi?ID=C12385136&Type=JANAFG&Table=on), Cp, is 20.79 J/mol-K. Therefore,
T2/T1 = (P2/P1)[1-1/γ]
T2/1200 = (1/20)[1-1/1.667]
T2 = 362 K
(h2-h1) = Cp(T2-T1)
(h2-h1) = 20.79*(362-1200)
(h2-h1) = -17420 J/mol
So each mole of atomic hydrogen has 17.42 kJ less energy at the turbine outlet than it had at the turbine inlet.
I used atomic hydrogen in this example because its heat capacity (or specific heat) and specific heat ratio is constant over all temperatures. The heat capacity and specific heat ratio of molecular gases vary with temperature (here's an example (http://www.engineeringtoolbox.com/hydrogen-d_976.html)), thus the calculations are more complicated. The calculations also include a mixture of different gases, and the composition of those gases vary with temperature. That is, the ratio of, say, H to H2 changes as the temperature increases or decreases, with the molecular form dominating at low temperature and the atomic form at high temperature.
Because the calculations can become insanely complicated, a computer program is typically needed or else it would take forever to do just one calculation.
This is the e-book for Sutton and Biblarz, the standard work. The thermodynamics chapter is skimmable without having to buy the book.
http://books.google.com/books?id=pFktw0GYSX8C
Thanks for the reference. Although I haven't had a chance to study it yet, I'm sure I'll find it interesting.
Another e-book I've found helpful is this one
http://books.google.com/books?id=TKdIbLX51NQC&printsec=frontcover#v=onepage&q&f=false
The part about gas generators starts on page 116. Note that page 118 states that the European Ariane uses one of the methods proposed by ka9q, i.e. using water as a diluent in a near-stoichiometric gas generator.
Also note that the gas generator example shown in Table 4-3 (page 119) indicates a LOX/RP-1 mixture ratio of 0.342, which is considerably lower than I was using. Redoing my calculations at the lower mixture ratio gives inlet and outlet temperatures of 1179 K and 891 K, and an enthalpy change of 1.2788 MJ/kg.
I've also redone my calculations using hydrogen peroxide at 90% concentration rather than 100% (It's my understanding 100% is unattainable). This changes the inlet and outlet temperatures to 1033 K and 524 K, with an enthalpy change of 0.8525 MJ/kg.
Yes, this is an awesome thread, and I'm sure LunarOrbit wouldn't object to splitting it off into the Reality forum.
I second that if Lunar Orbit doesn't mind doing it. We'd have to split off anything having to do with gas generators, preburners, and staged combustion.
Bob, what program are you using? Is it something I can rewrite to give you control over the ammonia dissociation?
I'm using freeware called STANJAN, written by Stanford University Mechanical Engineering Professor Bill Reynolds. The version I'm using has a 1984 copywrite. Being that old it's not very user friendly. I also had to create/format for myself most of the JANAF (http://kinetics.nist.gov/janaf/) data tables used by the program. Nonetheless, STANJAN has served me well over the years and I wouldn't have been able to accomplish half of what I have without it.
I've often wished I could write a Windows version that's easier to interface with, but I can't read the programming to figure out how it works. Just a few minutes ago a stumbled upon the following document, which might give me what I need to know. However, it's probably more of a pain in the neck to write a new program than to continue using what I already have.
http://www.stanford.edu/~cantwell/AA283_Course_Material/STANJAN_write-up_by_Bill_Reynolds.pdf
Regarding the ammonia dissociation issue, I apparently figured out how to force STANJAN to give me what I want because I found old notes from an investigation I did several years ago regarding monopropellant engines. At that time I calculated the temperature, molecular weight, and specific heat ratio of hydrazine at different percentages of dissociation. For the life of me I don't remember how I did it. I'll probably eventually figure it out again.
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(h2-h1) = Cp(T2-T1)
(h2-h1) = 20.79*(362-1200)
(h2-h1) = -17420 J/mol
Wait a minute, is that right? Cp is constant pressure heat capacity, but I'm not at constant pressure. Can anyone verify whether or not I goofed here? Nonetheless, the error, if it exists, is just in my one-time manual example calculation. I was using STANJAN for all the other calculations, so hopefully those are correct.
EDIT:
I think I might be OK. I plugged my example into STANJAN and I got a enthalpy change of -17270 J/g. Multiply that by the atomic weight of hydrogen and I get, -17270 x 1.008 = -17408 J/mol. I think that's close enough to my answer for verification. Either my calculation is correct, or both STANJAN and I are wrong.
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Wait a minute, is that right? Cp is constant pressure heat capacity, but I'm not at constant pressure. Can anyone verify whether or not I goofed here?
IIRC, the natural logarithm of the pressure ratio applies here somewhere, but I'd have to go back to references to determine where.
I'm using freeware called STANJAN, written by Stanford University Mechanical Engineering Professor Bill Reynolds.
[...]
I've often wished I could write a Windows version that's easier to interface with, but I can't read the programming to figure out how it works.
It's in Fortran 77.
[backs away slowly, not making eye contact]
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It's in Fortran 77.
I might be able to decipher how it works if I could view the source code; unfortunately all I have is an executable. Does anybody have a suggestion on how I might be able to decompile it?
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Fortran is easy enough to work with or decipher, but getting the source may be difficult -- the source originally had to be purchased (a Windows executable was free), but it appears that it is no longer supported (http://www2.galcit.caltech.edu/EDL/public/codes.html).
A newer, open-source package called Cantera (http://code.google.com/p/cantera/) may be worrh investigating.
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Wow, that brings back memories. I haven't touched FORTRAN for ages... and then it was the old FORTRAN IV, we never moved on to modern stuff like 77! ;D
Could I ask a favour - could one of you chaps recommend a beginners guide to rocket engines, if there is such a thing? I'd love to read a bit more on this if someone can point me in the right direction.
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It also (in the case of LOX/RP1 rockets, anyway) uses the stuff you already have large, mass-efficient tanks full of and which you already need to pump around
Yes, but...
Turbopumps in large rocket engines are powerful. As in far-more-powerful-than-a-speeding-locomotive powerful. A few messages back I computed a net mechanical power of 17.5 megawatts for the pumps in each F-1 engine. A typical Diesel locomotive generates only 2.2 - 2.5 megawatts. The turbopumps in the five F-1 engines on a S-IC stage required more power (87.5 MW) than all six Diesel engines on the late Costa Concordia combined (75.6 MW, total weight 1,000 tons).
Even the 100% efficient generation of that kind of power for several minutes requires a seriously large amount of fuel (and oxidizer). The less efficient the generation, the more fuel and oxidizer you'll need. Providing a separate fuel and/or oxidizer supply for the turbines might well be worthwhile from a mass standpoint if the turbine can operate more efficiently than it could by burning the main propellants.
Bob B. commented that staged cycle engines can achieve higher Isps by operating their main combustion chambers at higher pressures than gas-generator cycle engines. The F-1 engine (which uses the gas-generator cycle) would seem to bear this out; it achieves a sea level Isp of only 263 sec as compared to the RD-180, which uses a staged combustion cycle and achieves 311 sec at sea level. That's a big improvement.
Edited to add: Checking the F-1 documentation, I see that the actual turbine output power was 41 MW. I got my figure of only 17.5 MW from the actual work done by the pumps on the propellants; I knew the actual turbine power had to be considerably greater to overcome the usual losses. This further strengthens the case for more efficient turbines even at the expense of a dedicated fuel/oxidizer supply.
Edited again to add: Hey Bob, wanna figure the performance of a Saturn V in which the F-1 engines have been replaced with ten RD-180s?
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Bob B. commented that staged cycle engines can achieve higher Isps by operating their main combustion chambers at higher pressures than gas-generator cycle engines. The F-1 engine (which uses the gas-generator cycle) would seem to bear this out; it achieves a sea level Isp of only 263 sec as compared to the RD-180, which uses a staged combustion cycle and achieves 311 sec at sea level. That's a big improvement.
Staged combustion engines definitely provide better performance than gas generator engines, but they're also more complex and expensive to manufacture. When the engine gets thrown away after a few minutes of work, the cost has to be factored in.
The United States has favored cheaper gas generators, making up for the lower Isp by using high-performing LOX/LH2 in the core/upper stages. The Soviets/Russians never mastered LOX/LH2 propulsion. Instead they perfected staged combustion, thus getting better performance out of the engines rather than the propellant.
As far as I know, the only American staged combustion engine to reach full operational status was the SSME. Of course the SSME was reuseable, so the extra complexity and cost was deemed worthwhile. In expendable applications, the trend is still toward gas generators. In fact, the philosophy behind the Rocketdyne RS-68 engine was to build an engine as cheaply as possible at the sacrifice of some performance (gas generator, ablative cooling, etc.).
Cheaper still are solid rocket motors. I can't think of any Russian rockets that use SRMs, but they've been used widely in the USA for the last four decades. SRMs have lower Isp than liquid propellant engines, but they can deliever the same impulse for less money. They can also deliver very high thrusts, making them ideal for thrust augmentation at liftoff. Strapping a couple SRMs onto a liquid core can produce a high liftoff thrust-to-weight ratio, meaning a smaller percentage of the thrust is used in counteracting gravity. This makes the effective performance of solids a bit better than the Isp alone might imply.
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Staged combustion engines definitely provide better performance than gas generator engines, but they're also more complex and expensive to manufacture. When the engine gets thrown away after a few minutes of work, the cost has to be factored in.
I've been thinking -- instead of trying to make a vehicle like the shuttle that is completely reusable, why not make one where you retrieve and return only the expensive components like the high-performance upper stage engines, assuming you use cheap solids in the lower stages?
Somehow they would have to detach from the upper stage once in orbit, deorbit, protect themselves with a heat shield on the way down, and then land on parachutes.
I have no idea if this idea is remotely feasible, but it would be a compromise between fully expendable and fully reusable launchers, neither if which are getting any cheaper.