OK, I'm ducking for cover as I write this...but I figure that sometimes it's good to ask a really dumb question : the answers can be surprisingly enlightening...

Why do (most) traditional high-power liquid fuelled rocket engines have to inject the propellants into a combustion chamber at high pressure? That is, why do they need to use big turbopumps to get pressures of (I believe) 10 - 200 bar....why not just use smaller pumps to suck the fuel out of the tanks and inject it at lower pressure into the engine...if the rate of fuel use is the same, and if combustion efficiency is the same (itself debatable), why wouldn't this produce as high a thrust?

I say "most" engines because some, smaller engines are - I believe - just pressure fed, i.e. used Helium gas under pressure to force the propellants out of their tanks and into the engines. Were the LM engines like this? They seem to have respectable ISPs of ~300s.

I can think of these speculative reasons for the necessity of using high pressures :

-if you don't shoot the raw propellants in at high pressure, the force of the combustion happening in the chamber will prevent you from squirting any more in

-it lowers the volume of the combustion chamber (perhaps to an extent where you can get a lighter engine...the reduced chamber diameter could dominate over the need for thicker chamber walls in a higher pressure unit)

-with cryogenic propellants, if you don't keep them under high pressure they would have boiled off as they neared the hot engine. (Doesn't explain why you still use high pressure engines with non-cryogenic propellants)

-you get a contribution to the overall thrust from the fact that the propellants were injected under pressure (must be tiny though).

Any ideas which of the above - if any - dominate? Or are there other reasons?

Has anyone ever done calculations to see if you could make a large booster without turbopumps...using a pressure-fed system with lower ISP but without the weight, cost and complexity of the pumps and high-pressure combustion chamber? Do engines which just use (He) gas from a reservoir achieve input pressures close to those of turbopump-based units?

The higher the combustion chamber pressure, the higher the engine thrust and efficiency. The pressure on the downstream side of the pumps/upstream side of the injector must be greater than the combustion chamber pressure to force the propellant into the chamber. Although the pumps required to do this can be large and heavy, the added efficiency gained by a high chamber pressure reduces the amount of propellant that has to be burned to gain the same thrust. The amount of propellant saved more than offsets the weight of the pumps such that the overall weight of the rocket is less.

Pressure-fed systems are generally only feasible on a small scale because the entire propellant system must be pressurized, including the storage tanks. This means the thickness of the tank walls must be heavy enough to withstand the higher pressures, which results in heavy tanks. In a pump-fed system the tank walls can be very thin and lightweight because internal tank pressure is low. The only high-pressure part of the system is downstream of the pumps.

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Right, I follow. So the first answer I suggested was the correct one...you need to inject at high pressure, else you can't overcome the high pressure of the combusting propellants in the chamber and can't get any more propellants in.

And I see what you are saying about the consequences for the tanks, etc. if you have a gas-pressure fed design.

I suppose my conceptual troubles come in part from the fact that rocket equations (like Tsiolkovsky) use the exhaust velocity, and this is usually given in basic texts as being dependent only on the propellants, i.e. determined by chemistry not engine design. Therefore, in my mind, if you burn X and Y you get a reaction whose products have an energy determined by the chemistry alone, and this relates to a certain value of kinetic energy and hence exhaust velocity; and this will happen at whatever pressure you burn X and Y.

I guess I have not been able to reconcile this small-scale molecular view with the bulk view of the engine. In this view, it is clear that a higher combustion chamber pressure will result in gases emerging with higher exhaust velocity, as is the case for a jet engine or gun, etc. (Although in a rocket engine, you have had to expend some energy achieving those high pressures).

So I have one model of an engine where it is clear that you want a high combustion pressure (and you have explained why pumps are the way to get that for a large engine); and one model where it is the chemistry alone which matters in achieving a high Ve. What I lack is a way to fill in the gaps between those two views!

I say "most" engines because some, smaller engines are - I believe - just pressure fed, i.e. used Helium gas under pressure to force the propellants out of their tanks and into the engines. Were the LM engines like this? They seem to have respectable ISPs of ~300s.

As previously answered, the propellant pressure has to be greater than the chamber pressure or the propellant can't enter the combustion chamber. Some rockets have used pressurized tanks despite the inherent weight penalty. The Bell X-1 (technically, the XS-1) that Chuck Yeager flew to break the sound barrier used pressurized tanks. That plane had a maximum speed of about Mach 1.3. Later versions used turbopumps and took advantage of the lighter tanks to carry more propellant. The extra propellant (and some other improvements) allowed the X-1A to reach a top speed of Mach 2.435 (almost killing Yeager in the process).

The Apollo Service Module and Lunar Module engines used pressurized tanks with hypergolic propellants. Hypergolic propellants are nasty, toxic, and corrosive but they have the advantage of being storable for long periods and of igniting on contact. All they had to do on the SM and LM was to open the propellant valves and the engine would fire. This is much simplier and more reliable than firing up turbopumps and having to ignite the propellants. I don't know but suspect that NASA traded off heavier propellant tanks in exchange for simplicity and reliability.

Satellites commonly use pressurized hypergolic tanks for their propulsion systems for the same reason. I know of satellites that are still functioning over 15 years after launch and the storability of hypergolic propellants is a big contributing factor. You couldn't do that with just about any other type of propellant except probably ion engines.

Pressurized systems are cheaper but the weight penalty comes at a price. Turbopumps are expensive to develop and make but they are essential for large liquid fueled engines. Solid fueled and hybrid fueled rocket motors have the same weight penalty as pressurized tank systems because the engine casings have to be strong enough to withstand the combustion pressure. Sometimes, all it takes is a crack in the solid propellant to expose too much propellant to combustion, pushing the pressure past the casing rupture point. This is one of the more common causes of solid rocket motor failure.

This page might help if you want to learn some of the physics and math of rocket propulsion. Chamber pressure Pc is definitely a factor in exhaust gas velocity as equation 2.22 illustrates:



Furthermore, as the pressure goes up so does the combustion chamber temperature Tc. As you can see from the equation, Ve increases as Pc and Tc become larger.

Propellant chemistry is definitely a big part of the performance puzzle, but so is the design of the engine. This page has some charts that my help you see how different propellants perform relative to each other and relative to different chamber pressures. If you are not too adverse to math, you can try plugging different numbers into the above equation and see how the exhaust velocity is affected.

Another variable in the equation is the nozzle exit pressure Pe. Thrust is maximum when Pe is equal to the ambient air pressure, therefore a rocket's first stage engines will be designed with Pe close to one atmosphere and the upper stage engines will be designed with Pe close to zero. This is because the first stage engines must operate in the lower atmosphere and the upper engines operate in or near a vacuum.

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Thanks, people, for these seriously good answers.

The formulae and links look like exactly what I need - the simplistic idea that exhaust velocity is only dependent on the chemistry (or kinetics) of the reaction is clearly too simplistic (as I suspected), and this looks like a far more rigorous explanation.

Appreciate all comments about pressure-fed systems too. Makes me wonder though, since large solid-rocket motors clearly are practical and economic - despite not having the best Isp and despite the cost of the fuels and despite having to be thick and strong enough to contain the pressure - whether one could make a practical, economically viable, large liquid-fuelled booster which was pressure-fed. There may be no reason to do this of course, I accept that, unless good turbopump design was a technological entry barrier for some space efforts.

It's been suggested in the past, on the grounds that, while less efficient, it can be cheaper to built a big simple vehicle rather than a small complex vehicle. I recall a USAF "Big Dumb Booster" project in the 1960s and the OTRAG private venture.

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Also, Beale proposed to build big pressure-fed liquid fueled rockets. IIRC, he wanted to use kerosene and hydrogen peroxide as the propellants. To reduce the weight penalty, he was developing fibre wound composite tanks that would've been lighter than metal tanks of the same strength. Sadly, Beale pulled out of the booster business after spending several years and a lot of money but before attempting to launch. Beale blamed NASA but others claimed he made some lousy technical and practical decisions.

Happy to help. If you have any more questions, just ask.

Yes, and thanks for your fine explanation, Larry. In my previous post, in which I poo-pooed pressure-fed systems, I was referring to large launch vehicles where the majority of the vehicle weight is propellant and tanks. There are certainly applications, such as spacecraft maneuvering and attitude control systems, where pressure-fed systems are the way to go. In addition to their simplicity and reliability, pressure-fed systems can also be started and stopped an almost unlimited number of times as long as the propellant and pressurization gas holds out. Launch vehicle engines, with the possible exception of the upper most stage, need to start only once.

Partly offsetting this weight penalty it the fact that solid propellants are very dense, thus take up less volume. This means the casing can be smaller and lighter than the pressurized tanks required to hold an equivalent mass of liquid propellant.

I can think of a couple reasons why solid-propellant motors have an advantage over pressure-fed liquid systems. First is the density advantage, as already explained. Second, a solid propellant doesn’t need a pressurization system. The pressurization gas, the high-pressure tanks required to hold it, and the plumbing all adds mass that is unnecessary for a solid motor.

For large booster applications, solid-propellant motors seem to have an advantage over pressure-fed systems when simplicity is desired. And when you can cope with the complexity, a pump-fed system will give the best overall performance. This leaves the pressure-fed system as the odd man out.

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One of ways that that been tried to get around the need for pumps is the Roton, the rotary rocket. The idea was to create centrifugal force for the pumps by spinning the whole rocket. It never got off the ground, mostly due to funding problems by the looks, but there seems to be doubt over the basic design too.

Well, ya, a rotating rocket would have to be brought up to speed, somehow, before launch. There are a few solid rockets that actually do that - they are rotating when they are launched for stability.

Most cyrogenic liquid systems route the liquid fuel near the surface of the nozzle, preheating the fuel and cooling the nozzle.

The combustion chamber also has to make certain that there is intimate mixing of fuel and oxidizer. On early rockets, this was accomplished by forcing the gas through a peghole template - I don't know if this is still used. In general the bigger the motor, the more problematic fuel mixing becomes. Large solid rocket fuel/oxidizers are intimately mixed before ignition.

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jwj

It's ok not to know. We should try harder to find out.

The minimum pressure is about 2.7x ambient pressure for rockets to work at all, but the higher the better (mostly).
Yup. Then the pressure will fall, until the chamber pressure is below the propellant pressure, and then it will go in.
No, to a good approximation, they cancel out, for the same maximum thrust, the weight is roughly independant of maximum pressure. It's because a chamber is a pressure vessel; and the weight of pressure vessels are independent of the pressure of gas inside, but are proportional to the mass of gas that can be put inside.
That's nothing to do with it, you would just insulate.
Yes tiny.
No, the reason you want higher pressures is that you can (and usually have to) fit a bigger expansion ratio nozzle on the chamber. Bigger ratio nozzles convert more of the heat in the hot gas into kinetic energy, but they also reduce the pressure at the exit. For optimum efficiency you want the exit pressure to match ambient pressure, so that means you can't increase the expansion ratio too far; unless you increase the chamber pressure.

Hope this helps.

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Sea Dragon would self-pressurize on its own however.