JPH0653701U - Axial turbine blade structure - Google Patents

Abstract

(57) [Summary] [Object] To provide a structure of an axial flow turbine nozzle blade that can control the pressure distribution on the back surface of the blade and suppress the secondary flow generated between the blades from flowing into the mainstream side due to the pressure gradient. To do. [Structure] This is a structure of an axial flow turbine nozzle blade in which a line connecting the trailing edges of each cross section forms a straight line (L). The angle (θa) between the turbine axis and the line connecting the leading edge and the trailing edge in the central portion of the blade is the angle (θb) between the line connecting the leading edge and the trailing edge on the casing side and the hub side and the turbine axis. ,
It is set to be larger than θc).

Description

[Detailed description of the device]

[0001]

[Industrial applications]

 The present invention relates to a structure of an axial flow turbine nozzle blade used in a turbojet engine or a turbofan engine.

[0002]

[Prior art]

 FIG. 6 shows an example of a conventionally used gas turbine engine. In this gas turbine engine, the inflow air taken in through the air intake 1 is first compressed by the low pressure compressor 2. Then, a part of the compressed air sent from the low-pressure compressor 2 passes through the fan air discharge duct 3 and is directly injected outside the engine as a bypass thrust.

On the other hand, the compressed air sent from the low-pressure compressor 2 is further compressed at a high pressure by the high-pressure compressor 4, and then sent to the combustion chamber 5 where it is mixed with the supplied fuel, gasified and burned. The exhaust gas discharged from the combustion chamber 5 rotates the compressor turbine 6 and the fan turbine 7, and is ejected from the exhaust duct 8 as a core thrust. Both the compressor turbine 6 and the fan turbine 7 are composed of a row of stationary blades and a row of moving blades.

[0004]

[Problems to be solved by the device]

 By the way, in the gas turbine engine of this type, the shape of each blade 9 of the stationary blade row (nozzle) forming the compressor turbine 6 or the like is such that the gas discharged from the combustion chamber 5 causes the rear blade row to rotate. It is designed so that it can be converted into a suitable direction and speed, and is generally arranged diagonally as shown in FIG. 7, and the front edge is swollen and gradually becomes smaller toward the rear. Is formed in.

In the conventional turbine nozzle blade 9 of this type, when the flow of the combustion gas during operation is examined, a shear flow in the boundary layer occurs at a portion near the hub 10 and the casing 11, as shown in FIG. A secondary stream 12 is created. Then, there was a problem that the secondary flow 12 gradually spreads toward the downstream side and disturbs the flow of the main flow 13.

The present invention has been made in view of the above circumstances, and it is possible to control the pressure distribution on the back surface of a blade and suppress the secondary flow generated between the blades due to the pressure gradient from flowing into the mainstream side. An object is to provide a structure of a flow turbine nozzle blade.

[0007]

[Means for Solving the Problems]

 In order to achieve the above object, the present invention has an axial flow turbine nozzle blade structure in which the line connecting the trailing edges of each section is straight, and the line connecting the leading edge and the trailing edge in the central part and the turbine axis are The angle formed by is set larger than the angle formed by the turbine axis and the line connecting the leading edge and the trailing edge on the casing and hub sides.

[0008]

[Action]

 In the axial flow turbine nozzle blade according to the present invention, the angle formed by the blade back surface with the casing surface and the hub surface is an acute angle on the blade leading edge side. Therefore, due to the aerodynamic characteristics, the pressure in the portion close to the casing side and the hub side becomes smaller than the pressure in the central portion, and a pressure gradient from the central portion to the casing side or the hub side occurs. As a result, the secondary flow generated between the blades can be prevented from rolling up into the main flow (the central portion of the blade).

[0009]

【Example】

 Hereinafter, an axial flow turbine nozzle blade according to the present invention will be described with reference to FIGS. As shown in FIG. 1, the axial flow turbine nozzle blade 20 according to the present invention is arranged so as to be obliquely inclined with respect to the axial direction of the turbine (in other words, the combustion gas inflow direction), and the front edge is swollen. The diameter is gradually reduced toward the rear, and the entire surface is curved so that a concave portion is formed on the abdominal surface 20a and a convex portion is formed on the rear surface 20b.

A stacking method is one of the methods for determining the shape of the nozzle blade 20 of this type. This method is, for example, a method of determining each cross-sectional shape of the blade and stacking them in the turbine radial direction to determine the shape of the blade.

In the nozzle blade 20 shown here, the trailing edge 20 c is set to the stacking axis L. The stacking axis L is a straight line passing through the casing-side trailing edge Pa of the nozzle blade 20 and the hub-side trailing edge Pb, and the line connecting the trailing edges of the nozzle blade 20 in cross section is straight, that is, stacking. It will be located on the axis L. Further, in the nozzle blade 20, the angle θa formed by the line connecting the leading edge and the trailing edge on the mean (central portion) side of the nozzle blade and the turbine axis is the leading edge and the trailing edge on the casing side and the hub side. It is set to be larger than the angles θb and θc formed by the connecting line and the turbine axis.

Although the angles θa to θc vary depending on the size and output of the turbine, etc., they are generally set in the range of 35 ° to 65 ° as (θc, θb) <θa as shown in FIG. As specific values, for example, the angle θb on the casing side, the angle θa on the mean side, and the angle θc on the hub side are set to 55 °, 56 °, and 54 °, respectively. The angle of the cross-section of each blade located between these three cross-sections is set to a value that allows the angles of the blade cross-sections to be smoothly continuous (see Fig. 5).

Therefore, according to the axial flow turbine blade structure having the above-described configuration, a part of the gas flowing from the front combustor is divided into right and left at the front edge 20d of the nozzle blade 20. At this time, on the back surface 20b of the nozzle blade 20, a vortex due to the secondary flow 12 is generated near the casing 11 side and the hub 10 side due to the boundary layer at the nozzle blade inlet.

Here, the angle formed between the casing 11 and the hub 10 by the back surface 20b on the front edge 20d side of the nozzle blade is an acute angle. Therefore, due to the aerodynamic characteristics, the pressure on the mean side becomes larger than the pressure on the portion close to the casing side or the hub side, and a pressure gradient from the central portion to the casing side or the hub side occurs. As a result, the secondary flow generated between the blades can be prevented from rolling up into the main stream (the central portion of the blade), and the flow in the main stream can be prevented. Therefore, it is possible to supply a flow with less loss and a predetermined angle degree and less turbulence to the moving blades at the rear.

The structure of the axial turbine nozzle blade of the present invention is not limited to the above-mentioned embodiment, and the specific components such as the shape, material and size of each member can be appropriately changed in the implementation. is there. For example, in the above embodiment, the stacking axis L, which is a straight line passing through the casing-side trailing edge Pa of the nozzle blade 20 and the hub-side trailing edge Pb, passes through the center of rotation of the engine, but the stacking axis L is not necessarily the engine. Does not have to pass through the center of rotation, and may be inclined with respect to the inflow direction of the combustion gas.

[0016]

[Effect of device]

 According to the present invention, the line connecting the trailing edges of each cross section is straight, and the angle formed between the line connecting the leading edge and the trailing edge in the central portion of the blade and the turbine axis is the casing side of the blade and Since the angle is set to be larger than the angle between the turbine axis and the line connecting the leading edge and the trailing edge on the hub side, the pressure distribution on the back side of the nozzle blade is highest on the mean side and on the casing side and the hub side. Since the vortex due to the secondary flow generated between the blades does not wind up in the main flow part on the mean side, it does not disturb the flow of the gas flow and there is little loss and it is close to ideal It can have a wing shape.

[Brief description of drawings]

FIG. 1 is a schematic perspective view showing a structure of an axial flow turbine nozzle blade according to the present invention.

FIG. 2 is a plan view showing the structure of a coaxial flow turbine nozzle blade.

FIG. 3 is a view on arrow X in FIG.

FIG. 4 is a view on arrow Y in FIG.

FIG. 5 is a diagram illustrating an angle of each cross section of an axial turbine nozzle blade.

FIG. 6 is a schematic diagram showing the overall structure of a gas turbine engine.

FIG. 7 is a plan view showing a conventional example of an axial flow turbine nozzle blade.

[Explanation of symbols]

 20 axial flow turbine nozzle blade 20a axial flow turbine nozzle blade rear end 20b axial flow turbine nozzle blade front end

Claims (1)

[Scope of utility model registration request] 1. In the structure of an axial flow turbine nozzle blade in which a line connecting a trailing edge of each cross section is a straight line, an angle formed by a line connecting a leading edge and a trailing edge in a central portion and a turbine axis is on a casing side. And an axial flow turbine nozzle blade structure which is set to be larger than an angle formed by a line connecting the leading edge and the trailing edge on the hub side and a turbine axis.