Blade form for multi-stage axial compressor

Abstract

The invention relates to axial compressors of turbofan engines. The technical result is achieved by the use of a one-piece blade structure with a given geometry in a limited range of diameters D1 / D2, with the airfoil of the blade being described by arcs of circles, the geometric cross-sectional centres of which lie in a line along the length of the blade, thus providing for relative ease of manufacture, while the angle of inclination of the chord of the airfoil to the airflow, together with the chord length, the radii of a concave part of the airfoil section and the radii of a convex part of the airfoil section, provide for decent compressor wheel efficiency, while also providing sufficient strength and good pressure across a wide range of rotational speeds of the rotor of the compressor.

Description

[0001] Blade shape of a multistage axial compressor

[0002] The invention relates to axial compressors of fan gas turbine engines. Its application is most effective in axial compressors of fan turbojet low-power bypass engines with a thrust of 1000 to 5000 N. A fan gas turbine engine typically contains a compressor forcing an air flow into the engine combustion chamber. Compressors are of two types: axial and centrifugal. In an axial compressor, to create the necessary air flow pressure at the inlet to the combustion chamber, forcing is performed by passing the flow through a certain number of compressor wheels. The axial compressor wheel consists of a base on which a plurality of blades are fixed, forming a non-separable structure with the base, or blades fixed to the base with the possibility of removal.The need for low-power 1000-5000 N dual-circuit fan engines has led to the need to simplify the design of the engine compressor while maintaining its sufficiently high efficiency. The most critical for the efficiency of an axial compressor and labor-intensive to manufacture is its impeller with a large number of blades of a complex profile.

[0003] The design of a compressor blade is known (Invention Certificate RU No. 2150612 C1). The blade contains a section with a constant chord and a profile bending angle decreasing in height in accordance with the condition of constant theoretical pressure, as well as peripheral and root parts constituting 10-15% of the height of the latter with a profile bending monotonically increasing by 10-20% towards the ends in accordance with the condition of increased theoretical pressure. The peripheral and root parts of the blade are made with a profile chord monotonically increasing by 15-30% towards the ends. Such a design of the blade will ensure minimal profile losses.

[0004] The disadvantage of this design is the technological complexity of manufacturing the blade, as well as the concentration of stresses at the point where the profile transitions from straight to enlarged at the periphery.

[0005] A compressor blade design is known (Invention Certificate RU No. 2529272 C1). The axial compressor blade comprises an inlet edge, an outlet edge, a trough and a back with spherical vortex generators made on its surface, concave inside the blade. Each vortex generator is provided with at least two supply channels with outlet openings of diameter (0.05...0.25) D, equidistant from the vortex generator symmetry axis oriented in the direction of the oncoming flow. The distance from the vortex generator symmetry axis to the outlet opening is (0.1...0.4) D. The inlet openings of the supply channels are located on the trough of the blade, and the outlet openings are located at a distance of (0.025...0.7) D from the vortex generator leading edge. The feed channels are made at an angle of 20°...110° to the blade chord, where D is the diameter of the vortex generator imprint.The invention will increase the range of non-separating flow around the blades to 3%, increase the air flow through the compressor to 2% and increase the compressor efficiency to 4% by obtaining a stable vortex structure of the flow in vortex generators.

[0006] The disadvantage of the blade is the difficulty of producing a large number of through channels of small diameter, which results in the high cost of the impeller and the lack of prospects for use in low-power dual-circuit fan engines.

[0007] A known axial compressor blade (invention certificate RU No. 2516739 C1) is accepted as the closest analogue based on a number of features. The axial compressor blade contains a feather with a tail, has an S-shaped cross-section in the root part of the feather, to increase efficiency, the leading edge of the feather in the S-shaped area is bent toward the back of the blade, and the inflection point of the profile center line is located at a distance from the leading edge equal to 0.05-0.15 of the profile chord. The profile has good efficiency.

[0008] The disadvantage of the blade is the complexity of the profile shape, which limits the area of ​​application to high- and medium-thrust engines. Application in low-thrust engines is practically impossible due to the complexity of manufacturing blades or vanes of this profile.

[0009] The purpose of this invention is to develop a simple design of an axial compressor blade for low-power fan turbofan twin-circuit engines with a thrust of 1000 to 5000 N, with a relatively simple production technology, relatively low cost and comparatively high efficiency.

[0010] The set objectives are achieved by: using simple geometric shapes when constructing the blade profile, optimizing the blade chord angle to the air flow direction, optimizing the number of blades placed on the base of the compressor impeller for maximum performance and maximum efficiency of the compressor impeller. The geometric dimensions of the blade corresponding to the set objective are related by the following dependencies: the blade contains a convex approaching edge of radius R2 determined by the geometric construction of the profile, a concave trailing edge of radius R1, wherein the leading edge 1 and the trailing edge 2 of the blade are made straight and sharp, the root section of the blade profile 3 and the tip section of the blade profile 4 have the same geometry, the profile chord length L is constant its length is determined by the ratio L = k-7c D2 / N; where k is a coefficient taking into account the properties of the blade material from 1.65 to 2.1; л is a mathematical constant -3.1415;N is the number of compressor wheel blades, an even number in the range from 20 to 46, D2 is the inner diameter of the impeller, the angle of inclination of the profile to the flow axis A is in the range of 36-38°, the cone of narrowing of the impeller blade along the outer diameter B is in the range of 30-35°, the radius of curvature of the concave part is determined by the expression: R1=0 L; - flow property coefficient 2.8-3.1; the greatest blade thickness - t in is determined by the expression: t=L / y; where y is the coefficient taking into account the blade material 11.2-12.2; the radius of the convex surface of the blade profile, the ratio of the diameter of the outer impeller D1 to the inner D2 is determined by the volumetric flow rate of the pumped air, while the ratio D1 / D2 should not exceed 2.5.

[0011] The essence of the invention is explained by the drawings: Fig. 1 is an isometric view of the impeller of an axial compressor; Fig. 2 is the main view of the compressor wheel with the dimensions of the blade profile; Fig. 3 is a view of the impeller from the right.

[0012] Thus, the proposed technical solution for the compressor impeller blade profile allows for the manufacture of a solid impeller using widely used technological equipment, such as a CNC milling machine with five degrees of freedom, ensuring low production costs, while providing high rigidity of the compressor impeller structure, the rigidity of the impeller allows for the operation of the compressor rotor at high angular accelerations. The selected blade geometry and chord angle allow for maximum efficiency and high pressure under given initial conditions.

Claims

4 Invention formula Axial compressor blade of a fan gas turbine engine comprising a convex approach edge of radius R2 determined by the geometric design of the profile, a concave trailing edge of radius R1, wherein the leading edge 1 and the trailing blade 2 are made straight and sharp, the profile chord length L is constant, its length is determined by the ratio L = k n - D2 / N; where k is a coefficient taking into account the properties of the blade material from 1.65 to 2.1; y is a mathematical constant ~ 3.1415; N is the number of compressor wheel blades, an even number in the range from 20 to 46, D2 is the inner diameter of the impeller, the angle of inclination of the profile to the flow axis A is in the range of 36-38 °, the cone of narrowing of the blade wheel along the outer diameter B is in the range of 30-35 °, the radius of curvature of the concave part is determined by the expression: R1 = PL; is the coefficient of flow properties 2, 8-3.1; the greatest thickness of the blade - t in is determined by the expression: t = L / y; where y is the coefficient taking into account the blade material 11.2-12.2;the radius of the convex surface of the blade profile, the ratio of the diameter of the outer impeller D1 to the inner D2 is determined by the volumetric flow rate of the pumped air, while the ratio D1 / D2 should not exceed 2.5.;

Images