Why do gases in the combustion chamber only flow one direction to the gas turbine in a jet engine?

As far as I know from the working principle of jets engines, compressed air in the combustion chamber (or combustion canister) is mixed with fuel. The ignited mixture expands backwards to turn the turbine rotors and perpetuates the work cycle. Eventually, the left over hot gas creates the thrust.

Although heated gas expands in every direction in space, why does the combustion-gas only expand in one direction, towards the turbine?

• I asked that on Physics once. I'm not convinced I managed to explain to any of the responders what my question was. Jan 16, 2015 at 5:20
• Note that there are two ways to look at this: steady-state and engine start. If the engine wasn't running in the first place, you probably wouldn't be able to start it by injecting fuel and igniting that. You need an external mechanism to spool up the engine and start the gas flow before you start the combustion. Combustion can sustain, but not start the gas flow. Note that in flight, you can restart an engine using the forward speed of the plane to provide that gas flow. Jan 11, 2016 at 9:24
• If you assume combustion chamber is the highest pressure point in the engine then of course gas would want to flow both ways. But that's not true. The combustion chamber is the highest temperature but the compressors exit has the highest pressure, so the gas only flows towards to turbine side. Apr 30, 2020 at 15:38
• @albin that video is not available any more :/ Jul 4, 2022 at 12:31
• @HughPerkins you can find the working link at here youtu.be/gUA4xfmgSGI Jul 5, 2022 at 1:31

It is actually not that simple to ensure the proper gas speed in a gas turbine. In the compressor you want to limit the flow speed over the compressor vanes to the high subsonic range, so the inlet has to decelerate the flow down to approx. Mach 0.4 - 0.5. Less would mean less throughput and consequently less thrust.

This speed, however, is far too high for ignition. The fuel needs some time to mix with the compressed air, and if flow speed is high, your combustion chamber becomes very long and the engine becomes heavier than necessary. Therefore, the cross section leading from the compressor to the combustion chamber is carefully widened to slow down the airflow without separation (see the section below named "diffusor"). Around the fuel injectors you will find the lowest gas speed in the whole engine. Now the combustion heats the gas up, and makes it expand. The highest pressure in the whole engine is right at the last compressor stage - from there on pressure only drops the farther you progress. This ensures that no backflow into the compressor is possible. However, when the compressor stalls (this is quite like a wing stalling - the compressor vanes are little wings and have the same limitations), it cannot maintain the high pressure and you get reverse flow. This is called a surge.

The graph below shows typical values of flow speed, temperature and pressure in a jet engine. Getting these right is the task of the engine designer.

Plot of engine flow parameters over the length of a turbojet (picture taken from this publication)

The rear part of the engine must block the flow of the expanding gas less than the forward part to make sure it continues to flow in the right direction. By keeping the cross section of the combustor constant, the engine designer ensures that the expanding gas will accelerate, converting thermal energy to kinetic energy, without losing its pressure (the small pressure drop in the combustor is caused by friction and the Rayleigh effect). Now the accelerated flow hits the turbine, and the pressure of the gas drops in each of its stages, which again makes sure that no backflow occurs. The turbine has to take as much energy from the flow as needed to run the compressor and the attached pumps and generators without blocking the flow too much.

The remaining pressure is again converted to speed in the nozzle. Now the gas is still much hotter than ambient air, and even though the flow at the end of the nozzle is still subsonic in modern airliner engines, the actual flow speed is much higher than the flight speed. The speed difference between flight speed and the exit speed of the gas in the nozzle is what produces thrust.

Fighter engines usually have supersonic flow at the end of the nozzle, which requires careful shaping and adjustment of the nozzle contour. Read all about it here.

• "The rear part of the engine must block the flow of the expanding gas less than" less? Moire maybe?
– rbp
Jan 16, 2015 at 5:48
• @rbp: No, less is right. The flow always takes the path of least resistance, and if the turbine would block it more than the compressor, you get a surge. Jan 17, 2015 at 8:05
• @albin: I don't know myself. I googled the picture and think it shows the basics nicely. For a regular engine it has too little pressure remaining after the last turbine stage. It does not dive a big fan or a propeller, so there should be some energy left behind the turbine. The trend of the parameters looks right, however. Jan 18, 2015 at 7:03
• @PeterKämpf I have read that the combustion chambers are designed by considering the flow as Rayleigh flow. And in Rayleigh flow for subsonic case the static and stagnation pressures decrease during heat addition. So how does the pressure remain constant if we neglect the frictional loss. Apr 30, 2020 at 12:15
• @AbhishekPallipparagopakumar: Good remark, there is indeed a drop from the Rayleigh effect. Apr 30, 2020 at 13:29

The air in the compressor is both compressed and moving downstream towards the combustion section. The combustion does not create enough pressure to overcome all of that, and there is lower pressure as the air gets expanded through the turbine sections.

When the pressure in the compressor section drops too much, the combustion flames do expand in both directions. This is called a "compressor surge".

Disclaimer: I may have spent several hours on Wikipedia at one point trying to answer this question for myself!

Jet engines use the Brayton cycle, which is a "isobaric" process during combustion, meaning it keeps pressure constant during that phase. This is in contrast to the Otto cycle of a typical four-stroke piston engine, which is "isochoric" during combustion, meaning it keeps volume constant during that phase.

The Brayton cycle consists of 3 parts, of which combustion occurs in the middle

1. Incoming air is compressed. This requires work which is supplied by turbine at the end of the cycle while the engine is operating or an external motor when starting the engine. This raises the pressure of the air (decreasing the volume).
2. Fuel is mixed with the air and ignited. This is a continuous process (unlike piston engines). This process is "constant pressure," which is the part that you are asking about. It's not intuitively obvious why, so we'll dig into it in the next section. The end result is a larger volume of air at the same pressure as the inlet to the chamber from the compressor.
3. The air is passed through a turbine in a shape which decreases the pressure from the compressed pressure to atmospheric pressure. It uses this turbine to run the compressor. All pressure differential which is not needed to power the compressor is used to accelerate the air backwards to provide thrust.

So how does this "constant pressure" thing work? Treat the combustion chamber as a sort of box for a moment. Whether there is combustion or not, there's going to be generally constant pressure inside the box. Air is being pushed into it by the compressor, with some velocity and pressure. If the turbine at the end of the chamber can "dispose" of air at a high enough velocity, it can keep the pressure at the far end of the chamber equal to that of the front end.

So how does this "constant pressure" thing actually keep the flame front from advancing forward? The trick is that the flame front is trying to advance forward, but the velocity of the air through the chamber matches the speed of the flame front, keeping it in a constant location in the chamber. This is a dynamic process, so we'll need some dynamics. They key detail is that the turbine and compressor are on one axle, so what happens to one affects the other.

Consider three cases which compare the airflow velocity to the fuel:

• Too slow for combustion rate — this happens if you increase the throttle, or at the starting of the engine.
• The combustion starts to win, advancing the flame front towards the compressor, just like you'd think it would.
• However, now the dynamics come in. As the flame front moves forward, it raises the pressure at the outlet of the compressor, slowing the velocity of air through the compressor.
• This means the compressor doesn't have to pressurize as large of a volume of air, so its load decreases.
• (The compressor IS having to pressurize the volume to a higher pressure, but its trivial to see that this effect is overshadowed by the decreased volume of air by considering the edge case of a very weak compressor which clearly backs up if you pressurize it from the other side.)
• The turbine is still seeing the same pressure, but its less loaded. This spins the turbine up, increasing air flow leading us to...
• Trimmed engine — This is where planes normally try to operate, and it's where you don't see flames propagating forward.
• In this state, the velocity of the incoming air is enough to match the flame front's propagation forward.
• The flame front keeps trying to move forward, but air is shoved in front of it just as fast as it can burn.
• Too fast for combustion rate — this happens if you decrease the throttle.
• With less fuel, the flame front begins to fall backwards towards the turbine.
• As in the "too slow" case, dynamics come into play. As the flame front moves backwards, this also decreases the pressure on the inlet to the combustion chamber, and the compressor.
• The turbine now sees more load, spinning the turbine down. Slowing the turbine decreases airflow through the compressor, bringing us back towards a trim engine.

One takeaway from this pattern is why jet engines cannot change thrust quickly. If you rapidly add fuel by throttling up, you choke the engine, so you don't get much extra power until the turbine and compressor can spin up to equalize the engine.

• A flame front is only possible in a fuel/air mixture. The engine has fuel injectors only at the start of the combustor, where the flow speed is 20 - 40 m/s - way below any flame front speed. Ahead of this no fuel is present if the pressures are right. Actually, the combustor needs small flame holders to make sure that the flame remains in place! Jan 16, 2015 at 12:44
• "[...] turbine in a shape which decreases the pressure from the compressed pressure to atmospheric pressure." The turbine does not (necessarily) decrease the pressure down to atmospheric pressure - the atmospheric pressure is achieved by the nozzle after the turbine (although there are exceptions, see underexpanded and overexpanded nozzles).
– user12485
Dec 20, 2015 at 16:31
• @peterkampf I think you wrote that backwards. 20-40 m/s is way higher than the flame speed of any practical fuel.
– XRF
Apr 30, 2020 at 21:10
• @XRF: You're right, and I did not know that four years ago. That makes flame holders essential. Apr 30, 2020 at 22:48

I happened upon this question, and thought I could add some information, as I recognized the powerplant right away.

This is not another full answer, but only a response to @albin and @PeterKampf as to the engine model. I cannot yet add comments, so edit this as is necessary.

The engine picture you used in your answer is the power section of the Allison 501 series, which had several civil and military aircraft applications. The reduction gearbox and propeller was omitted from that picture (for whatever reason). That omission is confusing, please see my image below for the whole picture of what is happening in that powerplant. Usually an Aeroproducts or Hamilton-Standard propeller was fitted.

@PeterKampf, you were on the right track - there is little energy left over for jet thrust after the turbine has extracted it because most of the power needs to drive the RGB and propeller. Very typical of a turboprop (or turboshaft, for helicopters) design - residual jet thrust is of little use. In this case, given certain atmospheric conditions and at on-speed RPM (13,820), the turbine can extract roughly 10,000 horsepower from the gas stream. It takes ~6000 HP to drive the compressor, which leaves about 4000 shaft HP for the RGB and prop.

The Convair 580 uses 2 of these powerplants, The Lockheed C-130 Hercules uses 4.

If this were a turbojet design, the engine would likely have only 2 turbine stages instead of 4. This would extract only enough energy to drive the compressor (and thus self-sustaining the engine), with plenty of thrust left over for use.

• Welcome to aviation.SE. please be aware that this is not a discussion forum, but a Q&A site. I realize you want to comment on other answers, but this will only be available only after some positive contributions. Please stick around and you'll be able to comment under any post. Mar 13, 2019 at 13:04
• well, I for one found Bronco6363's response extremely useful. For example, I appreciated the details about how much power can be extracted from the turbine, how much is needed to drive the compressor (over half!), and so on. Jul 4, 2022 at 13:18