Limitations on gas turbine performance imposed by large turbine cooling flows

Calculations of the performance of modern gas turbines usually include allo wance 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. T here are two essential effects of turbine blade cooling: (i) the reduction of the gas stagnation temperature at exit from the combustion chamber (entr y to the first nozzle row) to a lower stagnation temperature at entry to th e first rotor and (ii) a pressure loss resulting from mixing the cooling ai r with the mainstream. Similar effects occur in the following cooled blade rows. The paper reviews established methods for determining the amount of c ooling air required and semi-empirical relations, for film cooled binding w ith thermal barrier coatings, are derived. Similarly, the pressure losses r elated to elements of cooling air leaving at various points round the blade surface are integrated over the whole blade. This gives another semi-empir ical expression, this time for the complete mixing pressure loss in the bla de row, as a function of the total cooling air used. These two relationship s are then used in comprehensive calculations of the performance of a simpl e open-cycle gas turbine, for varying combustion temperature and pressure r atio. These calculations suggest that for maximum plant efficiency there ma y be a limiting combustion temperature (below that which would tie set by s toichiometric combustion). For a given combustion temperature, the optimum pressure ratio is reduced by the effect of cooling air.