Improving the Cooling Air Supply System for the HPT Blades of High-Temperature GTE

Abstract

This paper describes the results of studies of the system for supplying cooling air to the HPT of high-temperature aviation bypass GTE. In the cooling cavity of the blade, a dividing partition is installed, which allows cold air to be supplied to the front cooling cavity of the blade, taken out of the high-pressure compressor, and to the rear cavity—air with lower pressure and temperature, taken from the intermediate stage of the compressor. Air cooled by the working blades of a GTE is fed into the tubes of a U-shaped air-to-air heat exchanger blown with air from the outer contour of this GTE. The results of the studies showed that the temperature of the air taken from the compressor in the AtA HE can be reduced by 110°–240°, depending on the geometric dimensions of the tubes and the configuration of the AtA HE. Problems to be solved: minimization of pressure losses in the external circuit of a gas turbine engine, development of methods for constructively increasing the intensity of air temperature reduction in tubular AtA HE and schemes for the optimal supply of this air to the inlet of cooled propeller blades. A tubular row-type AtA HE was designed, with micro-heat transfer intensifiers installed on the inner surface of small-sized thin-walled tubes, cylindrical or oval, into which cooled air drawn after the compressor or another, colder, but with lower pressure from its intermediate stage, is supplied. The system of cooling air cutoff, in the channels for supplying the rear cavity of the working blade of the turboprop engine on the cruising mode of GTE operation, implemented in the blades of the turbine rotor with a vortex matrix, is considered. In conclusion, the work presents recommendations on the design methodology of these units in modern and future aviation gas turbine engines.