JPH09256998A - Compressor wings - Google Patents

Abstract

(57) 【Abstract】 PROBLEM TO BE SOLVED: To reduce the rapid acceleration / deceleration of the flow at the leading edge of a blade to prevent the separation phenomenon of the boundary layer to improve the performance and to maintain the thickness of the leading edge to improve the durability. Keeps impact and wear resistance. SOLUTION: A front edge portion 1 of an air inflow side of a blade 12 of a compressor.
2a is formed in an asymmetrical shape centering on the mean camber line L so that the negative pressure surface side is thin and the positive pressure surface side is thick.

Description

Detailed Description of the Invention

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a compressor blade, and more particularly to a technique for increasing a compression efficiency by reducing an abrupt change in a flow at a leading edge portion.

[0002]

2. Description of the Related Art FIG. 4 shows a structural example of a gas turbine engine (turbo fan engine) used in an aircraft. In the figure, 1 is an air intake, 2 is a fan / low pressure compressor, 3 is a fan air discharge duct, 4 is a high pressure compressor,
5 is a combustion chamber, 6 is a high pressure turbine, 6a is a turbine shaft, 7
Is a low pressure turbine, 8 is an exhaust duct, 9 is a disk, 10
Is a moving blade, 11 is a casing, and 21 is a stationary blade.

In such an axial flow compressor portion as the fan / low pressure compressor 2 and the high pressure compressor 4 in the gas turbine engine, the rotor blade 10 is rotated by the disk 9 to compress the air to the rear side. I am sending it to.

[0004]

The blade 1 is provided between the disk 9 and the casing 11 (flow path) of the fan and both compressors.
The airflow is sent by the rotation of 0, and the concave side of the stationary blade 12 becomes the positive pressure surface 12B and the convex side of the stationary blade 12 becomes the negative pressure surface 12A.
However, at the leading edge, the flow hitting the blade tip is accelerated, and as shown by the airfoil Mach number shown in FIG. 6, a rapid acceleration / deceleration of the flow occurs. In a gas turbine engine compressor and other axial flow compressors, when a rapid deceleration of the flow occurs in the blade row, a boundary layer separation phenomenon is induced at the suction surface 12A of the stationary blade 12. However, there is a problem of reducing the compression efficiency. As a method for suppressing the rapid acceleration / deceleration of the flow at the leading edge of the blade, a method of reducing the thickness of the leading edge is generally adopted. However, if the leading edge is made thin, there remains a problem that impact resistance when foreign matter mixed in the air flow collides with the blade leading edge and abrasion resistance against sand, oxides, etc. in the air flow. .

The present invention has been made in view of these problems, and reduces the rapid acceleration / deceleration of the flow at the leading edge of the blade,
The purpose is to improve the performance by preventing the boundary layer from peeling, and to maintain the thickness of the leading edge of the blade to maintain impact resistance and wear resistance.

[0006]

The front edge of the compressor blade on the air inflow side is formed in an asymmetrical shape with the suction surface side thin and the pressure surface side thick with the mean camber line as the center. The length M of the asymmetric portion of the leading edge of the blade is set in the range of 0.01 <(M / C) <0.05 with respect to the cord length C of the blade. Further, in the asymmetrical portion of the leading edge portion of the blade, the thickness ta of the half blade portion on the suction surface side, which is located at a position M × 1/5 rearward from the front end, with the mean cabar line as a boundary, The thickness tb of the half blade on the pressure surface side is set in the range of 0.1 <(ta / tb) <0.9, and the radius of curvature Ra of the half blade on the suction surface side and the pressure surface side The radius of curvature Rb of the semi-wing portion has a relationship of Rb <Ra.

[0007]

[Function] When the front edge is formed in an asymmetric shape so that the suction surface side is thinner and the pressure surface side is thicker, a smoother air flow on the suction surface side can be secured as compared with the conventional symmetrical one. As a result, the rapid acceleration / deceleration of the flow at the front edge portion on the suction surface side is reduced. At this time, the thickness of the front edge portion of the blade becomes approximately the same as that of the conventional one by increasing the thickness of the positive pressure surface, and the impact resistance and wear resistance do not decrease. In this way, especially by reducing the rapid acceleration / deceleration of the flow on the suction side,
The occurrence of the back surface peeling phenomenon in the blade is suppressed, the thickness of the leading edge portion of the blade is maintained, and the impact resistance and wear resistance are maintained.

[0008]

BEST MODE FOR CARRYING OUT THE INVENTION Hereinafter, an embodiment in which a compressor blade according to the present invention is applied to a stationary blade of a compressor of a gas turbine engine will be described with reference to FIGS.

As shown in FIG. 1, the cross-sectional shape of the vane 12 is defined by a combination of transverse curves of an abdominal surface (positive pressure surface) 12B and a rear surface (negative pressure surface) 12A connecting the front edge portion 12a and the rear edge portion 12b. Is set. The leading edge portion 12a of the stationary blade 12 is
Conventionally, as shown by the chain double-dashed line in FIG. 2, generally, the negative pressure surface side and the positive pressure surface side have the same shape with the mean camber line L as the center. The front edge portion 12a is formed in an asymmetrical shape so that the suction surface side is thin and the pressure surface side is thick with the mean camber line L as the center.

Specifically, the length M of the asymmetric portion 20 of the leading edge portion 12a of the vane 12 is set so as to have the following relationship with the chord length C of the vane. 0.01 <(M / C) <0.05 where the ratio of the length M of the asymmetric portion to the cord length C is 0.
When it is 01 or less, as described later, it is not preferable in that the original function of making the asymmetric shape to reduce the rapid acceleration / deceleration of the flow at the front edge portion is diminished. When the ratio with the cord length C is set to 0.05 or more, the shape of the stationary blade 12 is significantly different from the initially designed shape, which is not preferable because the air compression performance is deteriorated. The total thickness t of the asymmetric portion 20 is almost the same as the conventional one. Therefore, the impact resistance when foreign matter mixed in the air flow collides with the blade leading edge portion and the wear resistance against sand, oxides, etc. in the air do not decrease.

Further, in the asymmetrical portion 20 of the front edge portion 12a of the stationary blade 12, on the negative pressure surface side with the mean cabar line L as a boundary, at a position which is recessed from the front end by a distance N of M × 1/5. The thickness ta of the half-blade portion 21A and the half-blade portion 2 on the pressure surface side
The thickness tb of 1B is set to have the following relationship. 0.1 <(ta / tb) <0.9 Here, the ratio of the thickness ta of the half-blade portion 21A on the suction surface side to the thickness tb of the half-blade portion 21B on the pressure surface side is set to 0.1 or less. If set, the tip shape of the leading edge portion 12a of the stationary blade 12 is greatly different from the initial design shape, and this is not preferable in that the air compression performance may be deteriorated. Conversely, on the contrary, the suction surface side half blade portion 21A Thickness of the semi-wing portion 21B on the pressure side and the thickness tb
It is not preferable that the ratio is set to 0.9 or more because there is no difference from the conventional one and the original function of reducing the peak Mach number at the leading edge part is not fulfilled.

The radius of curvature Ra of the half-blade portion 21A on the suction surface side of the asymmetric portion 20 and the radius of curvature Rb of the half-blade portion 12B on the pressure surface side are set to have the following relationship. Rb <Ra

FIG. 3 shows an example of analysis of the relationship between the chord length C of the stationary blade 12 and the Mach number of the blade surface in the embodiment under the above conditions. As is clear from this figure, the peak Mach number at the front edge portion on the suction surface side is greatly reduced compared to the conventional one. That is, the conventional peak Mach number before the shock wave on the suction surface side was about 1.1, but in the present embodiment, it decreases to about 0.97. In this way, since the peak Mach number at the leading edge is greatly reduced,
The change in speed before and after the leading edge is reduced, thereby reducing the pressure loss, and the air is compressed by the smooth and effective deceleration, which causes the separation phenomenon of the air flow on the suction surface (back surface) 12A of the stationary blade 12. It is estimated that the outbreak is suppressed.

In the above embodiment, the blade of the compressor according to the present invention is applied to the stationary blade of the compressor of the gas turbine engine, but the present invention is not limited to this. Needless to say, it can also be applied to moving blades.

[0015]

According to the blade of the axial flow compressor according to the present invention,
The following effects are obtained. (1) By forming the leading edge portion of the compressor blade on the air inflow side in an asymmetrical shape with the suction surface side thin and the pressure surface side thick with the mean camber line as the center, the blade at the leading edge portion is formed. The rapid acceleration / deceleration on the negative pressure surface is reduced while maintaining the thickness, and thus the compression efficiency can be improved without lowering the impact resistance and wear resistance. (2) By setting the length of the asymmetric part of the leading edge of the blade to the range of 0.01 <(M / C) <0.05 with respect to the cord length of the blade, the performance of the blade is reduced. It is possible to reduce sudden acceleration / deceleration at the front edge portion, prevent the boundary layer from peeling off on the back surface, and improve the compression efficiency to improve the performance. (3) Mx from the front end in the asymmetric part of the leading edge of the wing
The thickness of the half-blade on the suction surface side and the half-blade on the pressure surface side of the mean cabar line as a boundary at a position 1/5 behind is 0.1 <(ta / tb) <0.9, and the radius of curvature of the half-blade on the suction side and the radius of curvature of the half-blade on the pressure side are set to Rb <Ra. It is possible to further reduce the peak Mach number of the leading edge portion without deteriorating the performance, and thus it is possible to suppress application of excessive impact to the constituent material of the compressor.

[Brief description of drawings]

FIG. 1 is a cross-sectional view showing an embodiment of a stationary blade of an axial flow compressor according to the present invention.

FIG. 2 is an enlarged view of a portion X in FIG.

FIG. 3 is a relationship curve diagram between a chord direction position and a blade surface Mach number in the blade of FIG. 1;

FIG. 4 is a front sectional view showing a structural example of a gas turbine engine.

FIG. 5 is a cross-sectional view of a basic stationary vane in the example of FIG.

6 is a relationship curve diagram between a chord direction position and a blade surface Mach number in the stationary blade of FIG. 4;

[Explanation of symbols]

 2 fan / low pressure compressor 4 high pressure compressor 9 disk (hub) 10 moving blade 12A back surface (negative pressure surface) 12B belly surface (positive pressure surface) 20 asymmetric portion 21A negative pressure surface side half blade portion 21B positive pressure surface side half blade portion C Cord length

 ─────────────────────────────────────────────────── ─── Continuation of the front page (72) Kaoru Chiba, 229, Togaya, Mizuho-cho, Nishitama-gun, Tokyo Ishi Kawashima Harima Heavy Industries Ltd. Mizuho Plant Co., Ltd.

Claims (3)

[Claims] 1. A front edge portion (10a, 12a) of an air inflow side of a compressor blade (10, 12) is centered on a mean camber line (L) so that a suction surface side is thin and a pressure surface side is thick. A compressor blade characterized by having an asymmetrical shape. 2. The length (M) of the asymmetric portion (20) at the leading edge of the blade is 0.01 <(M / C) <0.05 with respect to the cord length (C) of the blade. A wing of a compressor, which is set in the range of. 3. In the asymmetric portion of the leading edge of the wing,
The thickness (ta) of the half-blade portion (12A) on the suction surface side and the half-blade portion on the pressure surface side, which are located rearward from the front end by a distance of M × 1/5, and which demarcates the mean cabar line. (21
The thickness (tb) of B) is set in the range of 0.1 <(ta / tb) <0.9, and the radius of curvature (Ra) of the half blade on the suction surface side and the half on the pressure surface side are set. The blade of the compressor according to claim 2, wherein the blade has a curvature radius (Rb) of Rb <Ra.