KR20240099892A - Turbine blade and gas turbine comprising it - Google Patents

Abstract

A turbine blade capable of improving durability by distributing torsional stress applied to the blade root member and a gas turbine including the same are disclosed.
The disclosed turbine blade is,
An airfoil with a leading edge and a trailing edge; A shank portion that secures the airfoil; A blade root member formed at the lower part of the shank and having dovetails on both sides; Stress that is formed on at least one of the leading edge corresponding portion corresponding to the leading edge and the trailing edge corresponding portion corresponding to the trailing edge among both sides of the blade root member in the axial direction to distribute the torsional stress applied to the blade root member. Includes dispersion grooves;

Description

Turbine blade and gas turbine comprising it}

The present invention relates to turbine blades and gas turbines including the same.

A turbine is a mechanical device that obtains rotational power through impulse or reaction force using the flow of compressible fluid such as steam or gas. There are steam turbines using steam and gas turbines using high-temperature combustion gas.

Among these, the gas turbine largely consists of a compressor, combustor, and turbine. The compressor is provided with an air inlet for introducing air, and a plurality of compressor vanes and compressor blades are alternately arranged within the compressor housing.

The combustor supplies fuel to the compressed air from the compressor and ignites it with a burner, thereby generating high-temperature, high-pressure combustion gas.

The turbine has a plurality of turbine vanes and turbine blades arranged alternately within the turbine casing. Additionally, the rotor is arranged to penetrate the center of the compressor, combustor, turbine, and exhaust chamber.

Both ends of the rotor are rotatably supported by bearings. Then, a plurality of disks are fixed to the rotor, each blade is connected, and a drive shaft such as a generator is connected to an end on the side of the exhaust chamber.

Since these gas turbines do not have a reciprocating mechanism such as the piston of a four-stroke engine, there is no mutual friction such as a piston-cylinder, so the consumption of lubricant is extremely low. The amplitude, which is a characteristic of a reciprocating machine, is greatly reduced, and high-speed movement is possible. There is an advantage.

To briefly explain the operation of a gas turbine, air compressed in a compressor is mixed with fuel and burned to produce high-temperature combustion gas, and the resulting combustion gas is injected toward the turbine. As the injected combustion gas passes through the turbine vanes and turbine blades, it generates rotational force, causing the rotor to rotate.

Republic of Korea Patent Publication No. 10-2010-0064754 (Name: Gas turbine cooling blade)

One aspect of the present invention is to provide a turbine blade that can improve durability by distributing torsional stress applied to the blade root member and a gas turbine including the same.

The turbine blade according to an embodiment of the present invention,

An airfoil with a leading edge and a trailing edge; A shank portion that secures the airfoil; A blade root member formed at the lower part of the shank and having dovetails on both sides; Stress that is formed on at least one of the leading edge corresponding portion corresponding to the leading edge and the trailing edge corresponding portion corresponding to the trailing edge among both sides of the blade root member in the axial direction to distribute the torsional stress applied to the blade root member. Includes a dispersion groove;

In the turbine blade according to an embodiment of the present invention, the dovetail is formed on both sides of the blade root member in the circumferential direction, a contact surface is formed on the outer surface in the radial direction, a plurality of dovetails are sequentially arranged along the radial direction, and a stress distribution groove is formed. Can be formed at a height corresponding to the dovetail disposed at the outermost radial direction in the blade root member.

In the turbine blade according to an embodiment of the present invention, the stress distribution groove includes a first stress distribution groove recessed in the leading edge corresponding portion with a first volume, and a second stress distribution groove recessed in the leading edge corresponding portion with a second volume. May include dispersion grooves.

In the turbine blade according to an embodiment of the present invention, the first volume may be formed to be larger than the second volume.

In the turbine blade according to an embodiment of the present invention, the stress distribution groove includes a first stress distribution groove whose width increases from the trailing edge side on the axial front to the leading edge side, and a trailing groove on the leading edge side on the axial back. It may include a second stress distribution groove whose width increases toward the edge.

In the turbine blade according to an embodiment of the present invention, the trailing edge side end of the first stress distribution groove has a first width, the leading edge side end has a second width greater than the first width, and the second stress distribution groove has a second width. The end on the leading edge side may have a third width, and the end on the trailing edge side may have a fourth width that is larger than the third width.

In the turbine blade according to an embodiment of the present invention, the second width of the first stress distribution groove may be formed to be larger than the fourth width of the second stress distribution groove.

In the turbine blade according to an embodiment of the present invention, the stress distribution groove includes a first stress distribution groove whose axial length gradually increases from the trailing edge side on the axial front to the leading edge side, and a leading edge on the axial back. It may include a second stress distribution groove whose axial length gradually increases from the side to the trailing edge side.

In the turbine blade according to an embodiment of the present invention, the first stress distribution groove increases in width from the trailing edge side on the axial front to the leading edge, and the second stress distribution groove has a trail on the leading edge side on the axial back. The width can increase toward the ring edge.

In the turbine blade according to an embodiment of the present invention, the stress dispersion groove is formed with a recessed depth of 1/20 to 1/40 of the axial length of the blade root member, and the recessed depth is the radial width of the stress dispersion groove. It can be formed from 1/2 to 2/5 of.

A gas turbine according to an embodiment of the present invention,

A compressor that takes in outside air and compresses it; A combustor that mixes fuel with air compressed in a compressor and combusts it; It includes a turbine in which turbine blades and turbine vanes are mounted inside, and the turbine blades rotate by combustion gas discharged from the combustor. Here, the turbine blade includes an airfoil having a leading edge and a trailing edge; A shank portion that secures the airfoil; A blade root member formed at the lower part of the shank and having dovetails on both sides; Stress that is formed on at least one of the leading edge corresponding portion corresponding to the leading edge and the trailing edge corresponding portion corresponding to the trailing edge among both sides of the blade root member in the axial direction to distribute the torsional stress applied to the blade root member. Includes dispersion grooves;

In the gas turbine according to an embodiment of the present invention, the dovetail is formed on both sides of the blade root member in the circumferential direction, a contact surface is formed on the outer surface in the radial direction, a plurality of dovetails are sequentially arranged along the radial direction, and a stress distribution groove is formed. Can be formed at a height corresponding to the dovetail disposed at the outermost radial direction in the blade root member.

In the gas turbine according to an embodiment of the present invention, the stress distribution groove includes a first stress distribution groove recessed in the leading edge corresponding portion with a first volume, and a second stress distribution groove recessed in the leading edge corresponding portion with a second volume. May include dispersion grooves.

In the gas turbine according to an embodiment of the present invention, the first volume may be formed to be larger than the second volume.

In the gas turbine according to an embodiment of the present invention, the stress distribution groove includes a first stress distribution groove whose width increases from the trailing edge side on the axial front to the leading edge side, and a trailing groove on the leading edge side on the axial back. It may include a second stress distribution groove whose width increases toward the edge.

In the gas turbine according to an embodiment of the present invention, the trailing edge side end of the first stress distribution groove has a first width, the leading edge side end has a second width greater than the first width, and the second stress distribution groove has a second width. The end on the leading edge side may have a third width, and the end on the trailing edge side may have a fourth width that is larger than the third width.

In the gas turbine according to an embodiment of the present invention, the second width of the first stress distribution groove may be formed to be larger than the fourth width of the second stress distribution groove.

In the gas turbine according to an embodiment of the present invention, the stress distribution groove includes a first stress distribution groove whose axial length gradually increases from the trailing edge side on the axial front to the leading edge side, and a leading edge on the axial back. It may include a second stress distribution groove whose axial length gradually increases from the side to the trailing edge side.

In the gas turbine according to an embodiment of the present invention, the first stress distribution groove increases in width from the trailing edge side of the axial front to the leading edge, and the second stress distribution groove has a trail on the leading edge side of the axial back. The width can increase toward the ring edge.

In the gas turbine according to an embodiment of the present invention, the stress dispersion groove is formed with a recessed depth of 1/20 to 1/40 of the axial length of the blade root member, and the recessed depth is the radial width of the stress dispersion groove. It can be formed from 1/2 to 2/5 of.

Details of other implementations of various aspects of the present invention are included in the detailed description below.

According to an embodiment of the present invention, durability can be improved by distributing torsional stress applied to the blade root member.

1 is a perspective view showing a partially cut away gas turbine according to an embodiment of the present invention.
Figure 2 is a cross-sectional view showing a schematic structure of a gas turbine according to an embodiment of the present invention.
Figure 3 is a partial cross-sectional view showing the internal structure of a gas turbine according to an embodiment of the present invention.
Figure 4 is a perspective view showing a turbine blade according to embodiments of the present invention.
Figure 5 is a diagram illustrating a turbine blade according to the first embodiment of the present invention as viewed from various sides.
Figure 6 is a diagram illustrating a turbine blade according to a second embodiment of the present invention as viewed from various sides.
Figure 7 is a diagram illustrating a turbine blade according to a third embodiment of the present invention as viewed from various sides.
Figure 8 is a side view of a turbine blade according to embodiments of the present invention viewed from the circumferential direction of the turbine.
9 and 10 are side views of modified examples of turbine blades according to embodiments of the present invention, viewed from the circumferential direction of the turbine.
Figure 11 is a side view of the blade root member in which stress dispersion grooves are formed in a straight line as viewed from the rotation axis direction of the turbine.

Since the present invention can be modified in various ways and can have various embodiments, specific embodiments will be exemplified and explained in detail in the detailed description. However, this is not intended to limit the present invention to specific embodiments, and should be understood to include all transformations, equivalents, and substitutes included in the spirit and technical scope of the present invention.

The terms used in the present invention are only used to describe specific embodiments and are not intended to limit the present invention. Singular expressions include plural expressions unless the context clearly dictates otherwise. In the present invention, terms such as 'include' or 'have' are intended to designate the presence of features, numbers, steps, operations, components, parts, or combinations thereof described in the specification, but are not intended to indicate the presence of one or more other features. It should be understood that this does not exclude in advance the possibility of the existence or addition of elements, numbers, steps, operations, components, parts, or combinations thereof.

Hereinafter, preferred embodiments of the present invention will be described in detail with reference to the attached drawings. At this time, in the attached drawings, identical components are indicated by identical symbols whenever possible. Additionally, detailed descriptions of well-known functions and configurations that may obscure the gist of the present invention will be omitted. For the same reason, some components are exaggerated, omitted, or schematically shown in the accompanying drawings.

Figure 1 is a perspective view showing a gas turbine according to an embodiment of the present invention partially cut away, Figure 2 is a cross-sectional view showing a schematic structure of a gas turbine according to an embodiment of the present invention, and Figure 3 is an implementation of the present invention. This is a partial cross-sectional view showing the internal structure of a gas turbine according to an example.

As shown in FIG. 1, a gas turbine 1000 according to an embodiment of the present invention includes a compressor 1100, a combustor 1200, and a turbine 1300. The compressor 1100 includes a plurality of blades 1110 installed radially. The compressor 1100 rotates the blade 1110, and the air moves while being compressed by the rotation of the blade 1110. The size and installation angle of the blade 1110 may vary depending on the installation location. In one embodiment, the compressor 1100 is directly or indirectly connected to the turbine 1300 to receive a portion of the power generated by the turbine 1300 and use it to rotate the blades 1110.

Air compressed in the compressor 1100 moves to the combustor 1200. The combustor 1200 includes a plurality of combustion chambers 1210 and a fuel nozzle module 1220 arranged in an annular shape.

As shown in FIG. 2, the gas turbine 1000 according to an embodiment of the present invention is provided with a housing 1010, and at the rear of the housing 1010 is a diffuser 1400 through which combustion gas passing through the turbine is discharged. ) is provided. Additionally, a combustor 1200 is disposed in front of the diffuser 1400 to receive compressed air and combust it.

If explained based on the direction of air flow, the compressor 1100 is located on the upstream side of the housing 1010, and the turbine 1300 is located on the downstream side. Additionally, a torque tube 1500 is disposed between the compressor 1100 and the turbine 1300 as a torque transmission member that transmits the rotational torque generated by the turbine 1300 to the compressor 1100.

The compressor 1100 is provided with a plurality of compressor rotor disks 1120 (for example, 14 pieces), and each of the compressor rotor disks 1120 is fastened by a tie rod 1600 so as not to be spaced apart in the axial direction. .

Specifically, each compressor rotor disk 1120 is aligned with each other along the axial direction with the tie rod 1600 constituting the rotation axis passing through approximately the center. Here, the opposing surfaces of each neighboring compressor rotor disk 1120 are compressed by the tie rod 1600, so that relative rotation is impossible.

A plurality of blades 1110 are radially coupled to the outer peripheral surface of the compressor rotor disk 1120. Each blade 1110 has a dovetail portion 1112 and is fastened to the compressor rotor disk 1120.

A vane (not shown) fixed to the housing is located between each rotor disk 1120. Unlike the rotor disk, the vanes are fixed so as not to rotate, and align the flow of compressed air passing through the blades 1110 of the compressor rotor disk 1120 to the blades 1110 of the rotor disk 1120 located on the downstream side. It plays a role in guiding the air.

The fastening method of the dovetail portion 1112 includes a tangential type and an axial type. It can be selected depending on the required structure of a commercial gas turbine, and may have a commonly known dovetail or fir-tree shape. In some cases, the blade may be fastened to the rotor disk using a fastening device other than the above type, for example, a fastener such as a key or bolt.

The tie rod 1600 is arranged to penetrate the centers of the plurality of compressor rotor disks 1120 and turbine rotor disks 1320, and the tie rod 1600 may be composed of one or more tie rods. One end of the tie rod 1600 is fastened to the compressor rotor disk located on the most upstream side, and the other end of the tie rod 1600 is fastened by a fixing nut 1450.

The shape of the tie rod 1600 may have various structures depending on the gas turbine, so it is not necessarily limited to the shape shown in FIG. 2. That is, as shown, one tie rod may have a shape that penetrates the central part of the rotor disk, or a plurality of tie rods may have a shape arranged circumferentially, and a combination of these may be possible.

Although not shown, in the compressor of a gas turbine, vanes that serve as guide blades may be installed at the next position of the diffuser to adjust the flow angle of the fluid entering the combustor inlet to the design flow angle after increasing the pressure of the fluid. This is called a deswirler.

The combustor 1200 mixes and combusts the incoming compressed air with fuel to produce high-energy, high-temperature, high-pressure combustion gas. The isobaric combustion process increases the temperature of the combustion gas to the heat resistance limit that the combustor and turbine parts can withstand. .

The combustors that make up the combustion system of a gas turbine may be arranged in large numbers in a housing formed in the form of a cell, including a burner including a fuel injection nozzle, a combustor liner forming a combustion chamber, and a combustor. It is composed of a transition piece that becomes the connection between the turbine and the turbine.

Specifically, the liner provides a combustion space where the fuel injected by the fuel nozzle is mixed with the compressed air of the compressor and burned. This liner may include a flame passage that provides a combustion space in which fuel mixed with air is combusted, and a flow sleeve that surrounds the flame passage and forms an annular space. Additionally, a fuel nozzle is coupled to the front end of the liner, and a spark plug is coupled to the side wall.

Meanwhile, at the rear end of the liner, a transition piece is connected to send combustion gas burned by the spark plug to the turbine side. The outer wall of this transition piece is cooled by compressed air supplied from the compressor to prevent damage due to the high temperature of combustion gas.

For this purpose, holes for cooling are provided in the transition piece to allow air to be sprayed inside, and the compressed air cools the body inside through the holes and then flows to the liner side.

Cooling air that cools the transition piece described above flows in the annular space of the liner, and compressed air from the outside of the flow sleeve may be provided as cooling air through cooling holes provided in the flow sleeve and collide with the outer wall of the liner.

Meanwhile, high-temperature, high-pressure combustion gas from the combustor is supplied to the turbine 1300. As the supplied high-temperature, high-pressure combustion gas expands, it collides with the rotor blades of the turbine, giving a recoil force to generate rotational torque. The rotational torque thus obtained is transmitted to the compressor through the torque tube 1500, providing the power necessary to drive the compressor. The excess power is used to drive generators, etc.

The turbine 1300 is basically similar to the structure of a compressor. That is, the turbine 1300 is also provided with a plurality of turbine rotor disks 1320 similar to the compressor rotor disk of a compressor. Accordingly, the turbine rotor disk 1320 also includes a plurality of turbine blades 2000 arranged radially. The turbine blade 2000 may also be coupled to the turbine rotor disk 1320 using a dovetail method or the like. In addition, turbine vanes 1330 fixed to the turbine casing 1350 are provided between the turbine blades 2000 to guide the flow direction of combustion gas passing through the turbine blades 2000.

The turbine vane 1330 is fixedly mounted within the turbine casing 1350 by a turbine vane platform 1340 coupled to the inner and outer ends of the turbine vane 1330. On the other hand, a ring segment 1360 is mounted at a position facing the outer end of the rotating turbine blade 2000 inside the turbine casing 1350 to form a predetermined gap with the outer end of the turbine blade 2000.

Hereinafter, turbine blades according to embodiments of the present invention will be described in detail with reference to FIG. 4. Figure 4 is a perspective view showing a turbine blade according to embodiments of the present invention.

Referring to FIG. 4, a turbine blade 2000 according to embodiments of the present invention includes an airfoil 2100, a shank portion 2200, a blade root member 2300, and a stress distribution groove 2400.

The airfoil 2100 is disposed outside the turbine blade 2000 in the turbine radial direction. Here, the turbine radial direction means the z-direction, hereinafter referred to as the radial direction (z-direction). The airfoil 2100 has an airfoil cross section and is formed to extend outward in the radial direction (z direction) of the turbine. A leading edge (LE) and a trailing edge (TE) are formed in the airfoil 2100. The leading edge (LE) is formed on the upstream side of the combustion gas flow, and the trailing edge (TE) is formed on the downstream side of the combustion gas flow.

A shank portion 2200 is disposed inside the radial direction (z direction) of the airfoil 2100. The shank portion 2200 may be formed in a substantially square plate shape. A cooling channel (not shown) through which cooling fluid can flow may be formed inside the airfoil 2100, and the cooling fluid may flow into the cooling channel through the shank portion 2200.

The blade root member 2300 is disposed inside the shank portion 2200 in the radial direction (z direction). For example, the inside of the shank portion 2200 in the radial direction (z direction) may be the lower portion of the shank portion 2200. The blade root member 2300 is formed to be narrow inward in the radial direction (z direction). Here, the width of the blade root member 2300 refers to the circumferential width, and the circumferential direction refers to the x-direction. Hereinafter, the x-direction is referred to as the circumferential direction.

Dovetails 2310, 2320, 2330, and 2340 are formed on both sides of the blade root member 2300 in the circumferential direction (x-direction). The dovetail has a fir-tree shape in cross section and can be formed in plural pieces. In this case, the dovetails 2310, 2320, 2330, and 2340 are a first dovetail 2310, a second dovetail 2320, a third dovetail 2330, and a fourth dovetail sequentially arranged inward in the radial direction (z direction). (2340) may be included.

From the first dovetail 2310 to the fourth dovetail 2340, the width of each dovetail 2310, 2320, 2330, and 2340 in the circumferential direction (x-direction) may gradually decrease. Here, the example of four dovetails is explained, but the number of dovetails is not limited to this.

The rotor disk 1320 is formed in a substantially disk shape. A plurality of grooves 1321 are formed on the outer periphery of the rotor disk 1320. The groove 1321 is formed to have a curved surface, and the blade root member 2300 of the turbine blade 2000 is inserted and coupled to the groove 1321. The dovetails 2310, 2320, 2330, and 2340 are inserted to engage with the groove 1321 of the rotor disk 1320. To this end, the cross-sectional shape of the groove 1321 of the rotor disk 1320 is formed to correspond to the cross-sectional shape of the dovetails 2310, 2320, 2330, and 2340 of the blade root member 2300.

When the rotor disk 1320 rotates, centrifugal force acts outward in the radial direction (z direction) on the blade root member 2300 inserted into the groove 1321, causing the dovetails 2310, 2320, 2330, and 2340. A contact surface (CS, see Figures 5 to 7 (a) and (b)) is formed in close contact with the groove of the rotor disk 1320. Since the direction of the centrifugal force acting on the blade root member 2300 is outside in the radial direction (z-direction), the contact surface CS is formed on the outside of the radial direction (z-direction) of the dovetail. In the accompanying drawings such as FIGS. 5 to 7, the contact surface CS is shown as being formed only on the first dovetail 2310, but this is only for convenience of explanation, and the second dovetail 2320 and the third dovetail 2330 ), a contact surface CS is also formed in each of the fourth dovetails 2340. Hereinafter, the contact surface CS refers to the contact surface CS of the first dovetail 2310.

The stress distribution groove 2400 is formed on at least one side of the blade root member 2300 on both sides in the axial direction of the turbine. Here, the axial direction means the y direction. The stress distribution groove 2400 is recessed into the inside of the blade root member 2300 and is formed to extend from the blade root member 2300 in the circumferential direction.

According to the analysis results of the torsional stress acting on the rotor disk 1320 in a state in which the stress distribution groove 2400 is formed in the blade root member 2300 in a straight shape as shown in FIG. 11, the blade without the stress distribution groove 2400 is formed. In the rotor disk 1320 with the root member 2300 assembled, the stress acting on the leading edge (LE) side was measured to be 1522 MPa, and the stress acting on the trailing edge (TE) side was measured to be 1632 MPa. In the case of the rotor disk 1320 assembled with the blade root member 2300 on which the stress distribution groove 2400 is formed, the stress acting on the leading edge (LE) side is 1202 MPa, and the stress acting on the trailing edge (TE) side is 1202 MPa. The stress was measured to be 1302 MPa.

That is, the stress acting on the rotor disk 1320 on which the blade root member 2300 with the stress distribution groove 2400 is assembled is the rotor on which the blade root member 2300 on which the stress distribution groove 2400 is not formed is assembled. It can be seen that the stress acting on the disk 1320 was reduced by 21.0% on the leading edge (LE) side and by 20.2% on the trailing edge (TE) side.

Meanwhile, as shown in FIG. 4, the airfoil 2100 may be arranged diagonally on the upper surface of the shank portion 2200. In this case, the torsional stresses F1 and F2 applied to the blade root member 2300 by the airfoil 2100 may be concentrated at both ends in the diagonal direction. That is, the first torsional stress F1 may be concentrated on a portion corresponding to the lower end of the leading edge LE (hereinafter also referred to as a “leading edge corresponding portion”). Additionally, the second torsional stress F2 may be concentrated on a portion corresponding to the lower end of the trailing edge TE (hereinafter also referred to as “trailing edge corresponding portion”).

Accordingly, in embodiments of the present invention, the stress distribution groove 2400 may be formed on both sides of the blade root member 2300 in the axial direction, at least one of the leading edge corresponding portion and the trailing edge corresponding portion. That is, the stress distribution groove 2400 is formed only on the leading edge (LE) side of the blade root member 2300, or on the leading edge (LE) side and the trailing edge (TE) side of the blade root member 2300. It can also form on both sides. The stress distribution grooves 2400 formed on both sides may be formed in a symmetrical or asymmetrical shape. The stress distribution groove 2400 may be formed at a height corresponding to the first dovetail 2310 disposed on the outermost side in the radial direction (z-direction) of the blade root member 2300.

Meanwhile, as combustion gas flows past the airfoil 2100, torsional stress occurs in the turbine blade 2000. Since the blade root member 2300 of the turbine blade 2000 is coupled to the rotor disk 1320, torsional stress is also applied to the dovetails 2310, 2320, 2330, and 2340 of the blade root member 2300 and the rotor disk 1320. It works. At this time, because the first dovetail 2310 is the dovetail located closest to the airfoil 2100 among the parts coupled to the rotor disk, torsional stress is most concentrated. Accordingly, the torsional stress can be distributed as the stress distribution groove 2400 is formed at a height corresponding to the first dovetail 2310.

A turbine blade according to a first embodiment of the present invention will be described with reference to FIG. 5. FIG. 5 is a view showing the turbine blade according to the first embodiment of the present invention as viewed from various sides, and FIG. 5 (d) shows the turbine blade 2000 according to the first embodiment of the present invention in the radial direction. (z-direction) is a plan view viewed from above, (a) in Figure 5 is a front view viewed from the front in the axial direction (y-direction) indicated by (d) in Figure 5, and (b) in Figure 5 is (d) in Figure 5 ) is a rear view seen from behind in the axial direction (y-direction) shown as B, and Figure 5(c) is a side view of Figure 5(d) as seen from the circumferential direction (x-direction) shown as C.

Referring to Figures 5 (a) to (d), the stress distribution grooves (2400: 2410, 2420) are formed by recessing the first volume in the leading edge (LE) corresponding portion where the first torsional stress (F1) is concentrated. 1 It may include a stress distribution groove 2410 and a second stress distribution groove 2420 recessed into a second volume in a corresponding portion of the trailing edge (TE) where the second torsional stress (F2) is concentrated. The shape of the volume forming the second stress distribution groove 2420 may be any shape.

Depending on the shape of the airfoil 2100 and the combustion gas flow, the first torsional stress F1 among the torsional stresses F1 and F2 may be greater than the second torsional stress F2. Therefore, the volume of the first stress distribution groove 2410 formed in the leading edge (LE) corresponding part is formed to be larger than the volume of the second stress distribution groove 2420 formed in the trailing edge (TE) corresponding part, so that stress is distributed. It is desirable to achieve equilibrium.

A turbine blade according to a second embodiment of the present invention will be described with reference to FIG. 6. FIG. 6 is a view showing the turbine blade according to the second embodiment of the present invention as viewed from various sides, and FIG. 6 (d) shows the turbine blade 2000 according to the second embodiment of the present invention in the radial direction. It is a plan view seen from above (z direction), (a) in Figure 6 is a front view seen from the front in the axial direction (y direction) indicated by (d) in Figure 6 (d) shown as A, and (b) in Figure 6 is (d) in Figure 6 ) is a rear view seen from behind in the axial direction (y-direction) shown as B, and Figure 6(c) is a side view of Figure 6(d) as seen from the circumferential direction (x-direction) shown as C.

Referring to (a) to (d) of FIG. 6, the stress distribution grooves 2400: 2410, 2420 increase in width from the trailing edge (TE) side of the axial (y direction) front side to the leading edge (LE) side. It may include a first stress distribution groove 2410 that increases in width, and a second stress distribution groove 2420 that increases in width from the leading edge (LE) side to the trailing edge (TE) side of the rear surface in the axial direction (y direction). there is.

That is, the first stress distribution groove 2410 is formed on the front surface in the axial direction (y direction) at a height corresponding to the first dovetail 2310 and has a length in the circumferential direction (x direction) corresponding to the first dovetail 2310, The end on the trailing edge (TE) side may have a first width (W1), and the end on the leading edge (LE) side may have a groove shape with a second width (W2) that is larger than the first width.

Similarly, the second stress distribution groove 2420 is formed on the axial (y-direction) rear surface at a height corresponding to the first dovetail 2310 and has a circumferential (x-direction) length corresponding to the first dovetail 2310. , the end on the leading edge (LE) side may have a third width (W3), and the end on the trailing edge (TE) side may have a groove shape with a fourth width (W4) that is larger than the third width.

In this way, the width of the leading edge (LE) corresponding portion where the first torsional stress (F1) is concentrated is increased on the axial (y-direction) front side, and the second torsional stress (F2) is concentrated on the axial (y-direction) back side. By increasing the width of the trailing edge (TE) corresponding portion, the durability of the blade root member 2300 can be improved by dispersing the torsional stress partially concentrated on the blade root member 2300.

Depending on the shape of the airfoil 2100 and the combustion gas flow, the first torsional stress F1 among the torsional stresses F1 and F2 may be greater than the second torsional stress F2. Therefore, the second width W2, which is the maximum width of the first stress distribution groove 2410, is formed to be larger than the fourth width W4, which is the maximum width of the second stress distribution groove 2420, to achieve equilibrium in stress distribution. desirable.

A turbine blade according to a third embodiment of the present invention will be described with reference to FIG. 7. FIG. 7 is a view illustrating the turbine blade according to the third embodiment of the present invention as viewed from various sides. FIG. 7 (d) shows the turbine blade 2000 according to the third embodiment of the present invention in the radial direction. It is a plan view seen from above (z direction), (a) in FIG. 7 is a front view looking at (d) in FIG. 7 from the front in the axial direction (y direction) shown as A, and (b) in FIG. 7 is (d) in FIG. 7 ) is a rear view seen from behind in the axial direction (y-direction) shown as B, and (c) in FIG. 7 is a side view of (d) in FIG. 7 as seen from the side in the circumferential direction (x-direction) shown as C.

Referring to (a) to (d) of FIG. 7, the stress distribution groove 2400 extends from the trailing edge (TE) side of the front axial direction (y direction) of the blade root member 2300 to the leading edge (LE) side. A first stress distribution groove 2410 that is extended and whose length in the axial direction (y direction) gradually increases from the trailing edge (TE) side of the front axial direction (y direction) to the leading edge (LE) side, and a blade root member It is formed to extend from the leading edge (LE) side of the axial (y-direction) back side of (2300) to the trailing edge (TE) side, and the trailing edge (TE) is formed from the leading edge (LE) side of the axial (y-direction) back side. It may include a second stress distribution groove 2420 whose length in the axial direction (y-direction) gradually increases toward the side.

That is, the first stress distribution groove 2410 is formed on the front surface in the axial direction (y direction) at a height corresponding to the first dovetail 2310 and has a length in the circumferential direction (x direction) corresponding to the first dovetail 2310, It may be in the form of a groove whose length in the axial direction (y-direction) gradually increases toward the leading edge (LE). The maximum length in the axial direction (y direction) of the leading edge LE may be the first length L1.

Similarly, the second stress distribution groove 2420 is formed on the axial (y-direction) rear surface at a height corresponding to the first dovetail 2310 and has a circumferential (x-direction) length corresponding to the first dovetail 2310. , It may be in the form of a groove whose axial (y-direction) length gradually increases toward the trailing edge (TE). The maximum length in the axial direction (y direction) at the trailing edge (TE) side may be the second length (L2).

In this way, the axial (y-direction) length of the leading edge (LE) corresponding portion where the first torsional stress (F1) is concentrated is increased on the axial (y-direction) front side, and the second torsional stress (F1) is increased on the axial (y-direction) back side. By increasing the axial (y-direction) length of the trailing edge (TE) corresponding portion where the stress (F2) is concentrated, the torsional stress partially concentrated on the blade root member 2300 is dispersed, thereby improving the durability of the blade root member 2300. can be improved.

Depending on the shape of the airfoil 2100 and the combustion gas flow, the first torsional stress F1 among the torsional stresses F1 and F2 may be greater than the second torsional stress F2. Therefore, the first length L1, which is the maximum length of the first stress distribution groove 2410, is formed to be larger than the second length L2, which is the maximum width and length of the second stress distribution groove 2420, so that the stress distribution It is desirable to achieve equilibrium.

The stress distribution groove 2400 of this embodiment can be merged with the second embodiment described above. That is, the width and length of the leading edge (LE) corresponding portion where the first torsional stress (F1) is concentrated are increased on the axial (y-direction) front side, and the second torsional stress (F2) is increased on the axial (y-direction) back side. Torsional stress can be distributed by increasing the width and length of the concentrated trailing edge (TE) counterpart.

Figure 8 is a side view of a turbine blade according to embodiments of the present invention viewed from the circumferential direction of the turbine.

The stress distribution groove 2400 may be formed by being recessed from the bottom of the shank portion 2200 in order to be formed from the portion of the blade root member 2300 closest to the airfoil 2100.

In particular, the torsional stress acts most intensively on the contact surface CS, which is the part where the rotor disk 1320 and the first dovetail 2310 are in close contact. In order to minimize the concentration of torsional stress, the stress distribution groove 2400 may be located in an area where the innermost depressed part of the blade root member 2300 has a height corresponding to the contact surface CS.

However, if the depression depth of the stress distribution groove 2400 into the blade root member 2300 is too large, the rigidity of the blade root member 2300 may be weakened. Therefore, the stress distribution groove 2400 must be recessed to a degree that can maintain the rigidity of the blade root member 2300 while distributing torsional stress.

According to the experimental and analysis results, in the embodiments of FIGS. 5 and 6, when the axial (y-direction) length of the blade root member 2300 is L and the recessed depth of the stress distribution groove 2400 is DP, When DP was greater than 1/20 of L, it was confirmed that the rigidity of the blade root member 2300 was rather reduced. In addition, it was confirmed that when DP is larger than 1/40 of L, the torsional stress dispersion effect is reduced. Accordingly, the recessed depth of the stress distribution groove 2400 may preferably be 1/20 to 1/40 of the axial (y-direction) length of the blade root member 2300.

In addition, in the embodiments of FIGS. 5 and 7, when the width of the stress distribution groove 2400 in the radial direction (z direction) is W, the recessed depth DP of the stress distribution groove 2400 is formed to be smaller than W. It can be. Preferably, DP may be formed from 2/5 to 1/2 of W. When formed in this way, the effect of torsional stress distribution can be improved.

Additionally, the stress distribution groove 2400 may include a first inclined surface (S1), a vertical surface (VS), and a second inclined surface (S2). The first inclined surface (S1), the vertical surface (VS), and the second inclined surface (S2) are parts sequentially arranged inward in the radial direction (z-direction) from the stress distribution groove 2400.

The first inclined surface (S1) is formed to be recessed into the inside of the blade root member 2300 as it goes inward in the radial direction (z direction) from the stress distribution groove 2400, and the second inclined surface (S2) is formed in the stress distribution groove (2400). It may be formed to be depressed into the inside of the blade root member 2300 as it goes outward in the radial direction (z-direction).

The vertical surface VS is formed between the first inclined surface S1 and the second inclined surface S2. The vertical surface (VS) is a portion where the surface is formed in a planar shape. The vertical surface (VS) is a portion whose cross-section is formed as a straight line perpendicular to the axial direction (y-direction) when the stress distribution groove 2400 is viewed from the circumferential direction (x-direction). In this case, the depth at which the stress distribution groove 2400 is recessed into the blade root member 2300 in the vertical plane (VS) can be kept constant along the radial direction (z-direction).

The portion depressed by the blade root member 2300 in the vertical surface VS may be the most deeply depressed portion among the portions depressed by the blade root member 2300 in the entire groove. In this case, the greatest dispersion of torsional stress can be formed in the vertical plane (VS).

Additionally, the vertical surface VS may be formed to overlap at least a portion of an area with a height corresponding to the contact surface CS of the first dovetail 2310. The radially (z-direction) inner portion of the vertical surface (VS) may overlap with the area of the contact surface (CS), and the radially (z-direction) outer portion of the vertical surface (VS) may overlap with the area of the contact surface (CS). In addition, the vertical surface VS may be included in the area of the contact surface CS, and the area of the contact surface CS may be included in the vertical surface VS. As described above, because torsional stress may be most concentrated in the contact surface CS, when the vertical surface VS is formed to overlap at least a portion of the area at the height corresponding to the contact surface CS, the distribution of torsional stress is It can be done effectively.

The first inclined surface S1, the vertical surface VS, and the second inclined surface S2 may be formed continuously with each other. That is, the first inclined surface S1, the vertical surface VS, and the second inclined surface S2 may have cross-sections that are continuous when the stress distribution groove 2400 is viewed in the circumferential direction (x-direction). When the first inclined surface (S1) and the second inclined surface (S2) are formed as a straight cross section viewed from the axial circumferential direction (x-direction) like the vertical surface (VS), the circumferential direction (x-direction) of the stress distribution groove 2400 The cross section seen from can be roughly formed as a trapezoid without a base. If the first inclined surface S1, the vertical surface VS, and the second inclined surface S2 are not formed in this way but are formed discontinuously, stress or load may be concentrated in the discontinuous portion.

9 and 10 are side views of modified examples of turbine blades according to embodiments of the present invention, viewed from the circumferential direction of the turbine.

According to the first modified example of the turbine blade shown in FIG. 9, the cross section of the first inclined surface S1 or the second inclined surface S2 of the stress distribution groove 2400 is formed as a curved surface. The cross section of the first inclined surface (S1) or the second inclined surface (S2) is preferably convex to the inside of the axial direction (y-direction) of the blade root member 2300 or to the outside of the axial direction (y-direction) of the blade root member 2300. It can be formed concavely.

The first inclined surface S1 and the second inclined surface S2 may be formed symmetrically with respect to the vertical surface VS. The first inclined surface S1, the vertical surface VS, and the second inclined surface S2 may be formed continuously with each other. In this case, the cross section of the stress distribution groove 2400 viewed from the circumferential direction (x-direction) is formed close to an arch shape, so that torsional stress can be distributed more effectively.

According to the second modification of the turbine blade shown in FIG. 10, the stress distribution groove 2400 has a dovetail 2310, 2320, 2330, 2340 in which the second inclined surface S2 is disposed at the outermost radial direction (z direction). It is formed to extend further inward in the radial direction (z direction).

In the turbine blades of the above-described embodiments, the stress distribution groove 2400 is formed only in the area of the height corresponding to the first dovetail 2310, but in this modified example, the stress distribution groove 2400 is formed at the first dovetail 2310. It may be formed by extending to other dovetails 2320, 2330, and 2340.

At this time, the vertical surface (VS), which is the most deeply recessed part of the stress distribution groove 2400, is disposed in an area at a height corresponding to the contact surface (CS) of the first dovetail 2310, and is located in a radial direction ( The second inclined surface S2 disposed on the inside (z direction) may be formed by extending to an area corresponding to the height of any one of the second dovetail 2320, the third dovetail 2330, and the fourth dovetail 2340. .

Above, an embodiment of the present invention has been described, but those skilled in the art can add, change, delete or add components without departing from the spirit of the present invention as set forth in the patent claims. The present invention may be modified and changed in various ways, and such modifications and changes will also be included within the scope of the rights of the present invention.

1000: Gas turbine
1100: Compressor
1200: Combustor
1300: Turbine
2000: Turbine Blades
2100: airfoil
2200: Shank part
2300: Blade root member
2400: Stress distribution groove
2410: first stress distribution groove
2420: Second stress distribution groove
VS: vertical plane
S1: first slope
S2: second slope

Claims (20)

An airfoil with a leading edge and a trailing edge;
A shank portion that secures the airfoil;
a blade root member formed below the shank portion and having dovetails on both sides;
Torsional stress applied to the blade root member by being formed on at least one of a leading edge corresponding portion corresponding to the leading edge and a trailing edge corresponding portion corresponding to the trailing edge among both sides of the blade root member in the axial direction. Stress distribution grooves that disperse;
Including turbine blades.
In claim 1,
The dovetails are formed on both sides of the blade root member in the circumferential direction, a contact surface is formed on the outer surface in the radial direction, and a plurality of dovetails are sequentially arranged along the radial direction,
The stress distribution groove is formed at a height corresponding to a dovetail disposed at the outermost radial direction in the blade root member.
The method of claim 1, wherein the stress distribution groove is,
A turbine blade comprising a first stress distribution groove recessed in a first volume in the leading edge counterpart, and a second stress distribution groove recessed in a second volume in the trailing edge counterpart.
In claim 3,
Turbine blade, wherein the first volume is formed to be larger than the second volume.
The method of claim 1, wherein the stress distribution groove is,
A turbine blade comprising a first stress distribution groove whose width increases from the trailing edge side of the axial front to the leading edge side, and a second stress distribution groove whose width increases from the leading edge side of the axial back side to the trailing edge side. .
In claim 5,
The trailing edge side end of the first stress distribution groove has a first width and the leading edge side end has a second width greater than the first width,
A turbine blade, wherein the leading edge side end of the second stress distribution groove has a third width, and the trailing edge side end has a fourth width greater than the third width.
In claim 6,
A turbine blade wherein the second width of the first stress distribution groove is formed to be larger than the fourth width of the second stress distribution groove.
The method of claim 1, wherein the stress distribution groove is,
A first stress distribution groove whose axial length gradually increases from the trailing edge side of the axial front to the leading edge, and a second stress distribution groove whose axial length gradually increases from the leading edge side of the axial back to the leading edge side. A turbine blade comprising dispersion grooves.
In claim 8,
The first stress distribution groove increases in width from the trailing edge side of the axial front surface to the leading edge side,
The turbine blade wherein the second stress distribution groove increases in width from the leading edge side of the axial rear surface to the trailing edge side.
The method according to any one of claims 1 to 9,
The stress dispersion groove is formed with a recessed depth of 1/20 to 1/40 of the axial length of the blade root member, and the recessed depth is 1/2 to 2/5 of the radial width of the stress dispersion groove. Formed as a turbine blade.
A compressor that takes in outside air and compresses it;
a combustor that mixes fuel with air compressed by the compressor and combusts it;
A turbine in which a turbine blade and a turbine vane are mounted, and the turbine blade rotates by combustion gas discharged from the combustor,
The turbine blade is,
An airfoil with a leading edge and a trailing edge;
A shank portion that secures the airfoil;
a blade root member formed below the shank portion and having dovetails on both sides;
Torsional stress applied to the blade root member by being formed on at least one of a leading edge corresponding portion corresponding to the leading edge and a trailing edge corresponding portion corresponding to the trailing edge among both sides of the blade root member in the axial direction. Stress distribution grooves that disperse;
Including gas turbines.
In claim 11,
The dovetails are formed on both sides of the blade root member in the circumferential direction, a contact surface is formed on the outer surface in the radial direction, and a plurality of dovetails are sequentially arranged along the radial direction,
The stress distribution groove is formed at a height corresponding to a dovetail disposed at the outermost radial direction in the blade root member.
The method of claim 11, wherein the stress distribution groove is,
A gas turbine comprising a first stress distribution groove recessed in a first volume in the leading edge counterpart, and a second stress distribution groove recessed in a second volume in the trailing edge counterpart.
In claim 13,
A gas turbine, wherein the first volume is formed to be larger than the second volume.
The method of claim 11, wherein the stress distribution groove is,
A gas turbine comprising a first stress distribution groove whose width increases from the trailing edge side of the axial front to the leading edge side, and a second stress distribution groove whose width increases from the leading edge side of the axial back side to the trailing edge side. .
In claim 15,
The trailing edge side end of the first stress distribution groove has a first width and the leading edge side end has a second width greater than the first width,
A gas turbine, wherein an end of the second stress distribution groove on the leading edge side has a third width, and an end on the trailing edge side has a fourth width that is larger than the third width.
In claim 16,
A gas turbine wherein the second width of the first stress distribution groove is formed to be larger than the fourth width of the second stress distribution groove.
The method of claim 11, wherein the stress distribution groove is,
A first stress distribution groove whose axial length gradually increases from the trailing edge side of the axial front to the leading edge, and a second stress distribution groove whose axial length gradually increases from the leading edge side of the axial back to the leading edge side. Gas turbine comprising dispersion grooves.
In claim 18,
The first stress distribution groove increases in width from the trailing edge side of the axial front surface to the leading edge side,
The second stress distribution groove is a gas turbine whose width increases from the leading edge side of the axial rear surface to the trailing edge side.
The method of any one of claims 11 to 19,
The stress dispersion groove is formed with a recessed depth of 1/20 to 1/40 of the axial length of the blade root member, and the recessed depth is 1/2 to 2/5 of the radial width of the stress dispersion groove. Formed as a gas turbine.

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