Limitations on gas turbine performance imposed by large turbine cooling flows

Calculations of the performance of modern gas turbines usually include allowance for cooling airflow rate; assumptions are made for the amount of the cooling air bled from the compressor, as a fraction of the mainstream flow, but this fractional figure is often set in relatively arbitrary fashion. There are two essential effects of turbine blade cooling: (i) the reduction of the gas stagnation temperature at exit from the combustion chamber (entry to the first nozzle row) to a lower stagnation temperature at entry to the first rotor and (ii) a pressure loss resulting from mixing the cooling air with the mainstream. Similar effects occur in the following cooled blade rows. The paper reviews established methods for determining the amount of cooling air required and semi-empirical relations, for film cooled binding with thermal barrier coatings, are derived. Similarly, the pressure losses related to elements of cooling air leaving at various points round the blade surface are integrated over the whole blade. This gives another semi-empirical expression, this time for the complete mixing pressure loss in the blade row, as a function of the total cooling air used. These two relationships are then used in comprehensive calculations of the performance of a simple open-cycle gas turbine, for varying combustion temperature and pressure ratio. These calculations suggest that for maximum plant efficiency there may be a limiting combustion temperature (below that which would tie set by stoichiometric combustion). For a given combustion temperature, the optimum pressure ratio is reduced by the effect of cooling air.