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6.2.4: Combustion chamber

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    The combustion chamber (also referred to as burner or combustor) is where combustion occurs. Fuel is mixed with the high-pressure air coming out of the compressor, and combustion occurs. The resulting high-temperature exhaust gas is used to turn the power turbine, producing the mechanical work to move the compressor and eventually producing thrust after passing through the nozzle.

    The burner is located between the compressor and the power turbine. The burner is arranged as some short of annulus so that the central engine shaft connecting turbine and compressor can be allocated in the hole. The three main types of combustors are annular; can; and hybrid can-annular.

    截屏2022-01-21 下午8.28.24.png
    Figure 6.7: Combustion chamber or combustor.

    Can combustors are self-contained cylindrical combustion chambers. Each can has its own fuel injector. Each can get an air source from individual opening. Like the can type combustor, can-annular combustors have discrete combustion zones contained in separate liners with their own fuel injectors. Unlike the can combustor, all the combustion zones share a common air casing. Annular combustors do not use separate combustion zones and simply have a continuous liner and casing in a ring (the annulus).

    Many modern burners incorporate annular designs, whereas the can design is older, but offers the flexibility of modular cans. The advantages of the can-annular burner design are that the individual cans are more easily designed and tested, and the casing is annular. All three designs are found in modern gas turbines.

    The details of mixing and burning the fuel are very complicated and therefore the equations that govern the combustion process will not be studied in this course. For the purposes of this course, the combustion chamber can be considered as the place where the air temperature is increased with a slight decrease in pressure. The pressure in the combustor can be considered nearly constant during burning. Using the station numbers from Figure 6.2, the combustor pressure ratio (CbPR) is equal to the stagnation pressure at stage 4 (\(p_{4t}\)) divided by the stagnation pressure at stage 3 (\(p_{3t}\)), i.e.:

    \[BPR = \dfrac{p_{4t}}{p_{3t}} \sim 1.\nonumber\]

    The thermodynamics in the combustion chamber are different from those of the compressor and turbine because in the combustion chamber heat is released during the combustion process. In the compressor and turbine, the processes are adiabatic (there is no heat involved): pressure and temperature are related, and the temperature change is determined by the energy equation.

    In the case of the combustion chamber, the process is not adiabatic anymore. Fuel is added in the chamber. The added mass of the fuel can be accounted by using a ratio \(f\) of fuel flow to air mass flow, which can be quantified as:

    \[f = \dfrac{\dot{m}_f}{\dot{m}} = \dfrac{\tfrac{T_{4t}}{T_{3T}} - 1}{\tfrac{\eta_b Q}{cT_{3t}} - \tfrac{T_{4t}}{T_{3T}}},\]

    where \(\dot{m}_f\) denotes the mass flow of fuel, \(Q\) is the heating constant (which depends on the fuel type), \(c\) represents the average specific heat, \(T_{t3}\) is the stagnation temperature at the combustor entrance, \(T_{4t}\) is the stagnation temperature at the combustor exit, and \(\eta_b\) is the combustor efficiency. This ratio is very important for determining overall aircraft performance because it provides a measure of the amount of fuel needed to burn a determined amount of air flow (at the conditions of pressure and temperature downstream the compressor) and subsequently generate the corresponding thrust.

    6.2.4: Combustion chamber is shared under a CC BY-SA 3.0 license and was authored, remixed, and/or curated by Manuel Soler Arnedo via source content that was edited to conform to the style and standards of the LibreTexts platform; a detailed edit history is available upon request.

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