# How do jet engines cool the gas in the combustion zone?

The adiabatic flame temperature of kerosene in air is 2093 C (source). Modern jet engines have a turbine inlet temperature (TIT) of 1700 to 1800 C. This is the max acceptable temperature before the turbine blades melt or weaken too much.

So how is the gas cooled to the lower temperature before it reaches the turbine stage?

Now I know the obvious simple answer is "air cooling". But I want to know some details about how it's achieved.

• Is the cooling air coming from the core air, or is it coming from some ducts into the combustion zone?

• Is the cooling air pumped by something (other than the compressor stages) near the core?

• Is the cooling air redirected turbulantly by some geometry inside the combustion zone?

• Is the cooling air expanded to achieve adiabatic cooling (again I'm only asking about inside the combustion zone)?

• Is fuel circulated around the core to act as a coolant (like rocket engines)?

When searching this, I came across something called inlet air cooling. This is not what I'm asking about. Inlet air cooling is about cooling the air before it enters the engine, for example by evaporative cooling.

Also, there is technology to cool the turbine blades directly, such as drilling holes in them and passing some other air . This is not what I'm asking about either. I want to know how combustion gas is cooled before it reaches the first turbine stage.

I realize of course that I can't get a full teaching of fluid dynamics from any answer on this site. What I'm hoping for at the very least is pictures, diagrams, and preferably some numbers used in certain equations that relate temperature and mass flow of air.

One more thing. Old engines are okay. In fact, they might be better, because older engines have a lower TIT and therefore must achieve even more cooling in the combustion zone. I did try to research them, e.g. the Rolls Royce Nene here, where a paragraph about cooling is summarized in a final sentence: "This diluting air is consequently expanded and accelerated rearwards, but it also cools the products of combustion to the temperature required at the turbine inlet." But it has very few numbers and no diagrams or cutouts.

• Blades and vanes can not tolerate temperatures anywhere near 1700 C. Oct 13, 2021 at 15:29
• From an older question you asked and I answered (and cited), the combustion doesn't combust all the air, but 22% of the air entering the combustor (give or take), "the rest will provide cooling, flame stabilization, and dilution."
– user14897
Oct 13, 2021 at 15:32

There's the fact that the hot gases from combustion are actually cooled in the combustion chamber. The turbine cooling isn't sufficient enough to maintain the appropriate temperature for the turbine blade(~1700 C).

The above figure(shamelessly taken from Aircraft Propulsion) clearly depicts the cooling mechanism in the combustion chamber(CC). 12% of air coming out of the compressor is primarily used for combustion and the rest of the air is used for cooling the hot gases. The temperature of the primary zone where the fuel is ignited can be greater than 2000 C. The cooling is done through film, transpiration, and impingement cooling. 20% of the air in the primary zone is used for better combustion efficiency and flame stability while (40%+20%=60%) of air in the dilution zone is used to maintain desired Turbine Inlet Temperature.

Is the cooling air coming from the core air, or is it coming from some ducts into the combustion zone?

• The core air is only used for cooling.

Is the cooling air pumped by something (other than the compressor stages) near the core?

• The cooling air is only pumped by the compressor.

Is the cooling air redirected turbulently by some geometry inside the combustion zone?

• The above figure clearly answers this question.

Is the cooling air expanded to achieve adiabatic cooling (again I'm only asking about inside the combustion zone)?

• As already mentioned, the cooling is done by transpiration, film, and impingement cooling.

Is fuel circulated around the core to act as a coolant (like rocket engines)?

• No, the fuel isn't used for cooling like in a rocket engine. It's inefficient as fuel isn't stored at cryogenic temperature.

Your source about the flame temperature is valid for a stoichiometric mixture. This means all oxygen molecules will be matched by carbon and hydrogen such that theoretically a complete conversion into H₂O and CO₂ can take place.

Real gas turbines, however, operate with a fairly large amount of excess air, which is why they produce comparably large amounts of nitrous oxides. This also means that energy from the combustion is not only heating the molecules involved in the combustion process, but also further oxygen molecules which will either remain or react with nitrogen. This excess oxygen is what allows afterburners to function - if all oxygen would be used up in the combustion chamber already, none would be left for the afterburner.

This excess air is used for dilution and as a cooling agent, and by being forced through holes or a porous structure in the single combustion cans or the liners of an annular combustor cools the structure defining the combustion chamber. To cite Wikipedia:

… combustors are carefully designed to first mix and ignite the air and fuel, and then mix in more air to complete the combustion process.

and further down, in the section about the liner:

… even though high performance alloys are used, the liners must be cooled with air flow.

The gas is not cooled prior to turbine inlet entry. Any attempt to cool down the gas, with decreasing efficiently of the engine. Most modern engines have turbine inlet temperatures of approximately 3600° F. This is far too hot for conventional alloys used to make the stator of rotor buckets to maintain acceptable yield strengths, so engineers have been forced to use the following techniques for manufacturing hot section components.

• Manufacturing parts from exotic ceramic materials or ceramic hybrids which offer the required material properties at these temperatures.

• Manufacturing parts form exotic metals such as nickel, cobalt, molybdenum, titanium or alloys of these.

• Exotic single crystal casting techniques which offer superior material properties to conventional manufacturing processes.

• Bleed air boundary layers - bleed air is tapped from the HPC and exits turbine casing, stators and rotors through small, laser drilled holes in the surface of the parts, effectively enveloping the part in a boundary layer of cooler air, preventing hot gas from making contact with the blades.

• The gas is not cooled prior to turbine inlet entry. This doesn't make sense to me. If the flame temp is 2093 C and the TIT is 1800 C, then obviously it was cooled. Any attempt to cool down the gas, with decreasing efficiently of the engine. Indeed, cooling the gas does decrease efficiency because we are creating a smaller difference in temperature in the Brayton Cycle. But it's necessary. The turbine can only handle 1800 C so there is no choice, we must cool it to 1800 C. Oct 13, 2021 at 16:45