Objective: Heat addition to a moving fluid.

Design Criteria:

a) Maximum efficiency of heat addition (add all of the available heat)
b) Minimum loss in stagnation pressure
c) Small length, volume, weight
d) Safety and reliability:

        i) Wall temperature must stay within material limits
        ii) Proper temperature distribution at the exit to control turbine blade stresses
        iii) No hot spots at walls ( no deposits )
        iv) No flame-out
        v) Easy to re-light if flame goes out.

Liquid Fuel is sprayed as a fine mist, mixed with air, and ignited. Note that heat is released only where both fuel and oxygen molecules are present in close proximity, in sufficient concentrations. There is an optimum ratio of fuel to air, called the stoichiometric ratio. The reaction occurs fastest where the mixture is stoichiometric. In the primary combustion zone, the mixture ratio should be close to stoichiometric, and the resulting temperature will be close to the adiabatic flame temperature of the fuel, considerably higher than the maximum allowable value of turbine inlet temperature. For example, note that the adiabatic flame temperature of hydrocarbon fuels can exceed 2500K, while the turbine inlet temperature is limited to about 2000K even in advanced military engines.

Near the walls of the combustor, colder air, rich in oxygen, is brought in, and a much "leaner" flame burns, with a lower temperature. The final mixture of burnt products and air will be considerably cooler than the primary zone.

Afterburners and Ramjet Combustors

More fuel can be added downstream of the turbine exit in gas turbine engines, and burned to reach higher temperatures than the turbine inlet temperature. This is usually done in a duct, where fuel is sprayed and the flame is stabilized using some means of generating low-speed recirculating flows. The fluid velocity and the Mach number are quite high in such burners, and the associated stagnation pressure losses are also quite high.

Combustor Pressure Losses

There are three main sources of stagnation pressure losses

1. Loss due to friction and flow separation (boundary layers, mixing devices, flameholders)
2. Shock losses due to deceleration of supersonic flows
3. Rayleigh Line losses due to heat addition at high Mach number.

In jet engine main combustors, Type 1 losses are the most important, because the Mach number is low. In afterburners and subsonic-combustion ramjets, Types 1 and 3 are important. All types are critically important in supersonic-combustion ramjets.

Rayleigh-Line Losses

1) Adding heat to a flowing fluid drives the Mach number towards 1.0.
2) Stagnation pressure loss is directly related to Mach number.
3) Thus, in a constant-area duct, with Station 1 denoting conditions upstream of heat addition and Station 2 the conditions downstream,


The decrease in stagnation pressure due to heat addition can be large.