Calculations of the performance of modern gas turbines usually include allowance for cooling air flow 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 blading 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 be set by stoichiometric combustion). For a given combustion temperature, the optimum pressure ratio is reduced by the effect of cooling air.

Averaging nonuniform flow is important for the analysis of measurements in turbomachinery and gas turbines; more recently an important need for averaging arises with results of computational fluid dynamics (CFD). In this paper we show that there is a method for averaging which is “correct,” in the sense of preserving the essential features of the nonuniform flow, but that the type of averaging which is appropriate depends on the application considered. The crucial feature is the decision to retain or conserve those quantities which are most important in the case considered. Examples are given to demonstrate the appropriate methods to average nonuniform flows in the accounting for turbomachinery blade row performance, production of thrust in a nozzle, and mass flow capacity in a choked turbine. It is also shown that the numerical differences for different types of averaging are, in many cases, remarkably small.

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A thermodynamic cycle analysis computer code for the performance prediction of cooled gas turbines has been used to calculate the efficiency of plants with varying combustor outlet temperature, compressor pressure ratio, and turbomachinery polytropic efficiency. It is shown that the polytropic efficiency exerts a major influence on the optimum operating point of cooled gas turbines: for moderate turbomachinery efficiency the search for enhanced combustor outlet temperature is shown to be logical, but for high turbomachinery efficiency this is not necessarily so. The sensitivity of the cycle efficiency to variation in the parameters determining the cooling flow rates is also examined. While increases in allowable blade metal temperature and film cooling effectiveness are more beneficial than improvements in other parameters, neither is as important as increase in turbomachinery aerodynamic efficiency.

In the early development of gas turbines, many empirical design rules were used; for example in obtaining fluid deflection using the deviation from blading angles, in the assumption of zero radial velocities (so-called radial equilibrium) and in expressions for clearance loss (the Lakshminarayana formulas). The validity of some of these rules, and the basic fluid mechanics behind them, is examined by use of modern ideas and computational fluid dynamics (CFD) codes. A current perspective of CFD in design is given, together with a view on future developments.