RU2311565C1 - Gas-turbine engine high-pressure compressor - Google Patents

Abstract

FIELD: mechanical engineering; compressors.

SUBSTANCE: according to invention, in high-pressure compressor of gas-turbine engine each stage of which includes working blade and guide vane installed after working blade, both installed in passage part, ratio of area F_in.1 of compressor passage part at inlet of first working blade to area F_out.1 of passage part at outlet from first stage is 1.6-2.2, ratio of area F_in.2 of compressor passage part at inlet of second working blade to area F_out.2 of passage part at outlet from compressor is 2.5-3.2 and Z - number of stages of compressor following the first stage is 4-6.

EFFECT: improved reliability and increased efficiency of compressor, reduced cost of manufacture and repair owing to increased gas dynamic stability and minimization of number of stages of rotary guide vanes.

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Description

The invention relates to high-pressure compressors of gas turbine engines, including for aviation applications.

Known axial two-stage compressor, in which gas-dynamic stability is ensured by rotation of the rotors of low and high pressure with different angular speeds [S.A. Vyunov. Design and engineering of aircraft gas turbine engines. Moscow, "Mechanical Engineering", 1981, p. 64, Fig. 3.8.b]. Such a compressor does not require special means of mechanization, however, its design drawback is reduced reliability due to the increased number of bearings, bearings and shafts.

Closest to the claimed is a high-pressure compressor of a gas turbine engine, in which the third, fourth, fifth, sixth and seventh stages are made high-pressure, i.e. with an increased degree of compression due to the "preload" of the flow part of the 3 ... 7 compressor stages [RF Patent No. 2243419, F04D 29/60, 2004]. The disadvantages of the known design are the low reliability characteristics and efficiency of the compressor due to the large number of stages, including the first two with rotary guide vanes.

The technical problem solved by the invention is to increase the reliability and efficiency of the compressor and reduce the cost of manufacturing and repair by increasing the reserves of gas-dynamic stability and minimizing the number of stages of rotary guide vanes.

The essence of the invention lies in the fact that in a high-pressure compressor of a gas turbine engine, each stage of which includes a working blade located in the flowing part and the following vanes of the guide apparatus, according to the invention, the ratio of the area F in. ₁ of the compressor running part at the entrance to the first working blade to the area F _chan.1 flow portion downstream of the first stage is 1.6-2.2, the ratio F _INP2 flow area of the compressor inlet to the second rotor blade to the area P _chan.2 flow part kompres ora output of the compressor is 2.5-3.2, and Z number of compressor stages following the first one, equal to 4-6.

The fulfillment of the ratios F _in.1 / F _out.1 = 1.6-2.2 and F _in.2 / F _out.2 = 2.5-3.2 allows us to produce the first stage of the compressor high-pressure and supersonic with the compression ratio π _{compression channel} ≈2,3 ... 2,4 (compression ratio π _k = 14) with the minimum number of stages (4-6). Subsequent to the first stage of the compressor perform low-pressure and subsonic. This embodiment can significantly reduce the number of rotary guide vanes in the compressor with a corresponding increase in reliability. At the same time, the time and costs for refining the compressor are significantly reduced due to the simplification of matching low-pressure stages with the high-pressure supersonic stage located in front of them.

Since a high-pressure stage is located at the compressor inlet, it is made with larger dimensions compared to subsequent stages and with thickened profiles, i.e. resistant to the dynamic effects of air flow and damage by foreign objects. This allows you to minimize the negative impact of spurious air leaks through the radial gaps between the stator and the rotor on the efficiency of the compressor.

Low-speed subsonic flow around subsequent to the first compressor stages allows to reduce hydraulic losses during flow around these stages with a corresponding increase in compressor efficiency as a whole, as well as thicken the profiles of the blades of these stages, which increases the reliability of the compressor. The use of the first high-pressure stage in the compressor can significantly reduce the cost of its manufacture due to the minimum number of stages and for the same reason reduce the cost of repairing the compressor.

The fulfillment of the ratio F _in1 / F _out.1 <1.6 leads to a decrease in the degree of compression in the first stage, which will require an increase in the degree of compression in the subsequent stages of the compressor with a corresponding decrease in the efficiency and gas-dynamic stability of the compressor.

With F _{input 1} / F _{output 1} > 2.2, hydraulic losses increase in the first supersonic compressor stage, which also leads to a decrease in compressor efficiency.

With F _in2 / F _out.2 <2.5, to preserve the overall degree of the compressor, a significant increase in the degree of compression in the first stage will be required.

An increase in hydraulic losses in the subsequent after the first stages of the compressor and a decrease in its efficiency is observed at F _{input 2} / F _{output 2} > 3.2.

At Z <4, the efficiency and reliability of the compressor will decrease, and at Z> 6, the number of parts and axial dimensions of the compressor will increase significantly.

The drawing shows a longitudinal section of a high-pressure compressor of a gas turbine engine of the claimed design.

The high-pressure compressor 1 of the gas turbine engine consists of a stator 2 and a rotor 3. In the flow part 4 between the stator 2 and the rotor 3 are placed, starting from the inlet 5 of the compressor 1, the blades 6 of the rotary inlet guide vane 7, the working blades of the first stage 8 and the rotary guide vanes 9 first stage. The working blades 8 and the guide blades 9 form the first high-pressure stage 10 of the compressor 1. Behind the first stage 10 there are working 11 and the guide 12 blades of the subsequent low-pressure stages 13 of the compressor 1.

The air stream 14 enters the input 5 of the compressor 1 and leaves the output 15.

The claimed device operates as follows.

When the high-pressure compressor 1 is operating, the air stream 14 from the inlet 5 enters a rotary input guide vane 7, from where it enters the working blades 8 of the high-pressure supersonic first stage 10. The kinetic energy of the air stream 14 accelerated by the blades 8 is converted into potential compression energy in the guide vanes 9 of the first stage 10.

Air 14 from the high-pressure first stage 10 enters the subsequent low-pressure stages 13, where it is further compressed with a high compression cycle efficiency.

Claims (1)

A high-pressure compressor of a gas turbine engine, each stage of which includes a working blade located in the flowing part and a guide _vane following it, characterized in that the ratio of the area F in. ₁ of the compressor running part at the inlet of the first working blade to the area F out. _{1 of the} flowing part at the outlet of the first stage is 1.6-2.2, the ratio of the area F _in.2 of the compressor flow _path at the entrance to the second working blade to the area F _out.2 of the compressor flow _path at the compressor outlet is 2.5-3.2 , but the number Z of compressor stages following the first is 4-6.