Turbine blade having non-axisymmetric endwall contour and gas turbine including the same

Abstract

To provide a turbine blade, and a gas turbine including the same, which are able to improve aerodynamic performance by reducing secondary vortices by forming a non-axisymmetric endwall contour having a plurality of convex portions and a plurality of concave portions from a rim seal at a leading-edge side.SOLUTION: A turbine blade includes: an airfoil including a pressure surface, a suction surface, a leading edge, and a trailing edge; an endwall formed integrally on a lower portion of the airfoil; and a root part formed integrally on a lower portion of the endwall, where an outer peripheral surface of the endwall is formed with a curved surface bent from a rim seal at one side to the other end.SELECTED DRAWING: Figure 5

Description

[Technical Field]

[0001] The present invention relates to a turbine blade having an asymmetric endwall contour and a gas turbine including the same. [Background technology]

[0002] A turbine is a mechanical device that uses the flow of compressible fluids such as steam or gas to generate rotational force through impulse or reaction force. Examples include steam turbines that use steam and gas turbines that use high-temperature combustion gases.

[0003] Among these, a gas turbine is broadly composed of a compressor, a combustor, and a turbine. The compressor is equipped with an air inlet for introducing air, and multiple compressor vanes and compressor blades are arranged alternately inside the compressor housing.

[0004] The combustor supplies fuel to the compressed air compressed by the compressor and ignites it with a burner, thereby generating high-temperature and high-pressure combustion gas.

[0005] The turbine has a plurality of turbine vanes and turbine blades arranged alternately within a turbine housing, and a rotor that passes through the center of the compressor, combustor, turbine, and exhaust chamber.

[0006] The rotor is rotatably supported at both ends by bearings. A plurality of disks are fixed to the rotor, and the blades are connected to each disk. At the same time, a drive shaft of a generator or the like is connected to the end of the rotor facing the exhaust chamber.

[0007] Such gas turbines do not have a reciprocating mechanism like the pistons in four-stroke engines, so there is no friction between the piston and cylinder, which means that lubricating oil consumption is extremely low. In addition, the amplitude that is characteristic of reciprocating machines is greatly reduced, allowing for high-speed operation.

[0008] In simple terms, the operation of a gas turbine is as follows: compressed air is mixed with fuel and burned in a compressor to produce high-temperature combustion gases, which are then injected into the turbine. The injected combustion gases pass through the turbine vanes and turbine blades, generating rotational force that rotates the rotor.

[0009] To construct such a turbine, a configuration is widely used in which a plurality of turbine rotor disks, each having a plurality of turbine blades arranged on its outer circumferential surface, are arranged in multiple stages so that the high-temperature, high-pressure combustion gas passes through the turbine blades.

[0010] Meanwhile, when turbine blades are assembled on-site, workers must assemble multiple turbine blades from the first stage turbine to the final stage turbine, which takes a lot of time and results in assembly tolerances after the work is completed.In addition, problems occur due to secondary vortexes when combustion gas moves because the steps or gaps between adjacent turbine blades are not specified.

[0011] When secondary vortices are generated when the combustion gas passes through the turbine blades, aerodynamic loss occurs on the intake surface or pressure surface.

[0012] In this case, the efficiency of the gas turbine decreases, and countermeasures are required to ensure stable movement of the combustion gas. [Prior art documents] [Patent documents]

[0013] [Patent Document 1] Republic of Korea Patent Publication No. 10-2019-0046118 (Published on May 7, 2019) Summary of the Invention [Problem to be solved by the invention]

[0014] An object of the present invention is to provide a turbine blade and a gas turbine including the same that can reduce secondary vortices and improve aerodynamic performance by forming an asymmetrical endwall contour having multiple protrusions and multiple recesses from the rim seal on the leading edge side. [Means for solving the problem]

[0015] In order to achieve the above object, the turbine blade of the present invention includes an airfoil including a pressure surface, a suction surface, a leading edge, and a trailing edge, an endwall integrally formed with a lower portion of the airfoil, and a root portion integrally formed with a lower portion of the endwall, and the outer peripheral surface of the endwall is formed as a curved surface that is bent from a rim seal on one side to the other end.

[0016] The end wall may include a first rim seal extending upstream from the leading edge and a second rim seal extending downstream from the trailing edge.

[0017] The outer peripheral surface of the end wall may be formed into a streamlined curved surface that connects from the first rim seal to the end portion on the trailing edge side.

[0018] The second rim seal may be positioned lower than the height of an outer peripheral surface of the end wall connected to the trailing edge.

[0019] The outer circumferential surface of the end wall may have a maximum radial height at a portion connected to the leading edge.

[0020] The radial height of the outer peripheral surface of the end wall may be increased from the first rim seal to a portion connected to the leading edge, and then decreased toward the trailing edge.

[0021] The outer peripheral surface of the end wall may include two recesses near the corners of the pressure face end between the sides of the leading edge and the trailing edge.

[0022] The outer peripheral surface of the end wall may include two recesses near the corners of the end on the suction surface side between the side of the leading edge and the side of the trailing edge.

[0023] The gas turbine of the present invention includes a compressor that takes in and compresses external air, a combustor that mixes fuel with the air compressed by the compressor and burns the resulting mixture, and a turbine in which turbine blades and turbine vanes are installed inside a turbine casing and the turbine blades are rotated by combustion gas discharged from the combustor. The turbine blade includes an airfoil having a pressure surface, an inlet surface, a leading edge, and a trailing edge, an endwall integrally formed with a lower portion of the airfoil, and a root portion integrally formed with a lower portion of the endwall, and the outer peripheral surface of the endwall is curved from a rim seal on one side to the other end.

[0024] The end wall may include a first rim seal extending upstream from the leading edge and a second rim seal extending downstream from the trailing edge.

[0025] The outer peripheral surface of the end wall may be formed into a streamlined curved surface that connects from the first rim seal to the end portion on the trailing edge side.

[0026] The second rim seal may be positioned lower than the height of an outer peripheral surface of the end wall connected to the trailing edge.

[0027] The outer circumferential surface of the end wall may have a maximum radial height at a portion connected to the leading edge.

[0028] The radial height of the outer peripheral surface of the end wall may be increased from the first rim seal to a portion connected to the leading edge, and then decreased toward the trailing edge.

[0029] The outer peripheral surface of the end wall may include two recesses near the corners of the pressure face end between the sides of the leading edge and the trailing edge.

[0030] The outer peripheral surface of the end wall may include two recesses near the corners of the end on the suction surface side between the side of the leading edge and the side of the trailing edge. [Effects of the Invention]

[0031] According to the turbine blade of the present invention and the gas turbine including the same, secondary vortices can be reduced and aerodynamic performance can be improved by forming an asymmetrical endwall contour having multiple protrusions and multiple recesses from the rim seal on the leading edge side. [Brief explanation of the drawings]

[0032] [Figure 1] 1 is a partially cutaway perspective view of a gas turbine according to an embodiment of the present invention; [Figure 2] 1 is a cross-sectional view showing a schematic structure of a gas turbine according to an embodiment of the present invention. [Figure 3] 1 is a partial cross-sectional view showing the internal structure of a gas turbine according to an embodiment of the present invention. [Figure 4A] FIG. 1 is a partial perspective view of a turbine blade according to the prior art. [Figure 4B] 1 is a partial perspective view showing a turbine blade according to an embodiment of the present invention. [Figure 5] 1 is a perspective view showing a turbine blade according to an embodiment of the present invention. [Figure 6] FIG. 6 is a perspective view of the turbine blade of FIG. 5 as viewed from the pressure surface side. [Figure 7] FIG. 6 is a perspective view of the turbine blade of FIG. 5 as viewed from the leading edge side. [Figure 8] FIG. 6 is a perspective view of the turbine blade of FIG. 5 as viewed from the suction surface side. [Figure 9] FIG. 2 is a top view showing the height of the outer peripheral surface of the end wall of a turbine blade according to an embodiment of the present invention, represented by contour lines. [Figure 10A] 1 is a photograph showing secondary vortices generated in a turbine blade according to the prior art. [Figure 10B] 4 is a photograph showing a secondary vortex generated in a turbine blade according to an embodiment of the present invention. [Figure 11] 4 is a graph showing pressure loss as a function of span position for a turbine blade according to the prior art and a turbine blade according to an embodiment of the present invention. DETAILED DESCRIPTION OF THE INVENTION

[0033] Although the present invention can be implemented in various forms by adding various modifications, specific embodiments will be illustrated and described in detail in the detailed description. However, it should be understood that this is not intended to limit the present invention to the specific embodiments, and that the present invention includes all modifications, equivalents, or alternatives that fall within the spirit and technical scope of the present invention.

[0034] The terms used in the present invention are merely used to describe specific embodiments and are not intended to limit the present invention. A singular expression includes a plural expression unless the context clearly indicates otherwise. It should be understood that, in the present invention, terms such as "comprise" or "have" are intended to specify the presence of features, numbers, steps, operations, components, parts, or combinations thereof described in the specification, and do not preclude the possibility of the presence or addition of one or more other features, numbers, steps, operations, components, parts, or combinations thereof.

[0035] Hereinafter, preferred embodiments of the present invention will be described in detail with reference to the accompanying drawings. It should be noted that in the accompanying drawings, identical components are denoted by the same reference numerals whenever possible. Detailed descriptions of known functions and configurations that may obscure the gist of the present invention will be omitted. For the same reasons, some components in the accompanying drawings may be exaggerated, omitted, or shown schematically.

[0036] FIG. 1 is a partially cutaway perspective view of a gas turbine according to one embodiment of the present invention, FIG. 2 is a cross-sectional view showing a schematic structure of a gas turbine according to one embodiment of the present invention, and FIG. 3 is a partially cross-sectional view showing an internal structure of a gas turbine according to one embodiment of the present invention.

[0037] As shown in FIG. 1 , a gas turbine 1000 according to one 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 arranged radially. The compressor 1100 rotates the blades 1110, and the rotation of the blades 1110 compresses and moves air. The size and installation angle of the blades 1110 can be changed depending on the installation position. In one embodiment, the compressor 1100 is directly or indirectly connected to the turbine 1300, and receives a portion of the power generated by the turbine 1300 and uses it to rotate the blades 1110.

[0038] The air compressed by the compressor 1100 travels to a combustor 1200. The combustor 1200 includes multiple combustion chambers 1210 and fuel nozzle modules 1220 arranged in an annular configuration.

[0039] 2, a gas turbine 1000 according to one embodiment of the present invention includes a housing 1010, and a diffuser 1400, through which combustion gas that has passed through the turbine is discharged, is provided on the rear side of the housing 1010. A combustor 1200, which receives and combusts compressed air, is disposed in front of the diffuser 1400.

[0040] In terms of the air flow direction, compressor section 1100 is located upstream of housing 1010, and turbine section 1300 is located downstream. Between compressor section 1100 and turbine section 1300, torque tube unit 1500 is located as a torque transmission member that transmits the rotational torque generated in turbine section 1300 to compressor section 1100.

[0041] The compressor section 1100 is provided with a plurality of (for example, 14) compressor rotor disks 1120, and the compressor rotor disks 1120 are fastened together by tie rods 1600 so as not to be spaced apart in the axial direction.

[0042] Specifically, the compressor rotor disks 1120 are aligned axially with a tie rod 1600, which constitutes a rotation axis, penetrating substantially the center of each disk. Opposing surfaces of adjacent compressor rotor disks 1120 are pressed together by the tie rod 1600, preventing relative rotation.

[0043] A plurality of blades 1110 are radially coupled to the outer circumferential surface of the compressor rotor disk 1120. Each blade 1110 has a dovetail portion 1112 and is fastened to the compressor rotor disk 1120.

[0044] A vane (not shown) is fixed to the housing between each rotor disk 1120. The vane is fixed so as not to rotate differently from the rotor disks, and serves to align the flow of compressed air that has passed through the blades of the compressor rotor disk and guide the air to the blades of the rotor disk located downstream.

[0045] The dovetail portion 1112 can be fastened in either a tangential or axial manner, depending on the required structure of the commercial gas turbine. The dovetail portion 1112 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-described types, such as a key or a bolt.

[0046] The tie rod 1600 is disposed to penetrate through the centers of the compressor rotor disks 1120 and the turbine rotor disk 1320, and may be composed of one or more tie rods. One end of the tie rod 1600 is fastened into the compressor rotor disk located most upstream, and the other end of the tie rod 1600 is fastened by a fixing nut 1450.

[0047] The shape of the tie rod 1600 may be various depending on the gas turbine, and is not necessarily limited to the shape shown in Fig. 2. That is, as shown in the figure, it may have a shape in which one tie rod passes through the center of the rotor disk, or a shape in which multiple tie rods are arranged circumferentially, or a combination of these may be used.

[0048] Although not shown, the compressor of the gas turbine may be provided with a vane that acts as a guide vane at the position next to the diffuser in order to adjust the flow angle of the fluid entering the combustor inlet after increasing the pressure of the fluid to the design flow angle; this is called a deswirler.

[0049] The combustor 1200 mixes the incoming compressed air with fuel and burns it to produce high-energy, high-temperature, high-pressure combustion gas, and the temperature of the combustion gas is raised to the heat limit that the combustor and turbine components can withstand through the constant pressure combustion process.

[0050] The combustors that make up the combustion system of a gas turbine may be arranged in multiple numbers within a housing formed in a cellular shape, and are composed of a burner including a fuel injection nozzle, a combustor liner that forms a combustion chamber, and a transition piece that serves as a connection between the combustor and the turbine.

[0051] Specifically, the liner provides a combustion space where fuel injected from a fuel nozzle is mixed with compressed air from the compressor and burned. The liner may include a flame tube that provides the combustion space where the fuel mixed with air is burned, and a flow sleeve that surrounds the flame tube to form an annular space. The liner also has a fuel nozzle coupled to its front end and a spark plug coupled to its side wall.

[0052] A transition piece is connected to the rear end of the liner so that combustion gases burned by the spark plug can be sent 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 the combustion gases.

[0053] For this purpose, the transition piece is provided with cooling holes so that air can be injected into the interior, and the compressed air passes through the holes to cool the main body inside before flowing toward the liner.

[0054] Cooling air that has cooled the transition piece flows through the annular space of the liner, and compressed air is provided outside the flow sleeve as cooling air through cooling holes provided in the flow sleeve and collides with the outer wall of the liner.

[0055] Meanwhile, the high-temperature, high-pressure combustion gas discharged from the combustor is supplied to the turbine 1300. As the high-temperature, high-pressure combustion gas expands, it collides with the turbine rotors, generating a reaction force and a rotational torque. The rotational torque thus obtained is transmitted to the compressor via the torque tube, and any power exceeding that required to drive the compressor is used to drive a generator, etc.

[0056] The turbine 1300 is basically similar in structure to a compressor. That is, the turbine 1300 also includes a plurality of turbine rotor disks 1320 similar to the compressor rotor disks of a compressor. Accordingly, the turbine rotor disks 1320 also include a plurality of turbine blades 1340 arranged radially. The turbine blades 1340 can also be connected to the turbine rotor disk 1320 by a method such as a dovetail. In addition, turbine vanes 1330 fixed to the housing are also provided between the blades 1340 of the turbine rotor disk 1320 to guide the flow direction of the combustion gas passing through the blades.

[0057] 3, the turbine vane 1330 is fixedly mounted within the housing by a vane carrier 1335, which is an endwall connected to the inner and outer ends of the turbine vane 1330. Meanwhile, a ring segmont 1345 is mounted at a position facing the outer end of the turbine blade 1340, which rotates inside the housing, so as to form a predetermined gap with the outer end of the turbine blade 1340. In other words, the gap between the ring segmont 1345 and the outer end of the turbine blade 1340 forms the tip clearance.

[0058] FIG. 4A is a partial perspective view showing a turbine blade according to the prior art, and FIG. 4B is a partial perspective view showing a turbine blade according to one embodiment of the present invention.

[0059] The turbine blade 10 according to the prior art shown in Figure 4A includes an airfoil 11 and an endwall 12 integrally formed with the lower part of the airfoil. The airfoil 11 includes a pressure surface, a suction surface, a leading edge, and a trailing edge. The outer peripheral surface of the endwall 12 is formed in a substantially flat shape. Strictly speaking, the outer peripheral surface of the endwall 12 is formed in a curved shape having a constant radius of curvature about the rotation axis of the turbine 1300.

[0060] 4B shows a turbine blade 100 according to an embodiment of the present invention, which includes an airfoil 110 and an endwall 120 integrally formed with the lower portion of the airfoil. The endwall 120 may have rim seals formed on the upstream and downstream sides in the axial direction. In the present invention, the outer circumferential surface of the endwall 120 may be smoothly curved, connected to the upper surface of the upstream rim seal.

[0061] FIG. 5 is a perspective view showing a turbine blade according to one embodiment of the present invention, FIG. 6 is a perspective view of the turbine blade of FIG. 5 viewed from the pressure surface side, FIG. 7 is a perspective view of the turbine blade of FIG. 5 viewed from the leading edge side, and FIG. 8 is a perspective view of the turbine blade of FIG. 5 viewed from the suction surface side.

[0062] A turbine blade 100 according to one embodiment of the present invention includes an airfoil 110 including a pressure surface 111, a suction surface 112, a leading edge 113, and a trailing edge 114, an endwall 120 integrally formed with the lower portion of the airfoil, and a root portion 130 integrally formed with the lower portion of the endwall.

[0063] The airfoil 110 includes a pressure surface 111 formed by a recess on one side, a suction surface 112 formed by a bulge on the other side, a leading edge 113 formed at the upstream corner, and a trailing edge 114 formed at the downstream end.

[0064] The end wall 120 can be integrally connected to the lower portion, i.e., the radially inner end, of the airfoil 110. The outer peripheral surface of the end wall 120 may be formed into a curved surface that curves from the rim seal on one side to the other end.

[0065] The end wall 120 may include a first rim seal 121 extending upstream from the leading edge 113 side and a second rim seal 122 extending downstream from the trailing edge 114 side.

[0066] A plurality of turbine blades 100 are mounted on the circumference of the turbine rotor disk 1320. A first rim seal 121 of the endwall 120 may be configured to seal against a fixed upstream turbine vane 1330, and a second rim seal 122 may be configured to seal against a fixed downstream turbine vane 1330.

[0067] The outer peripheral surface of the end wall 120 may be formed into a streamlined curved surface that connects from the first rim seal 121 to the end on the trailing edge 114 side.

[0068] As shown in FIG. 5, the upper surface of the end wall 120 may be formed into a smooth, streamlined curved surface that is continuous with the upper surface of the first rim seal 121 .

[0069] The second rim seal 122 may be positioned lower than the height of the outer circumferential surface of the endwall 120 that is connected to the trailing edge 114 .

[0070] The second rim seal 122 may be formed to extend radially inwardly downstream from a corner at the downstream end of the outer circumferential surface of the end wall 120 .

[0071] As shown in FIG. 6, the first rim seal 121 can be positioned slightly higher than the second rim seal 122, but the radial height of the first rim seal 121 may be positioned slightly lower than the corner of the downstream end of the outer peripheral surface of the end wall 120.

[0072] As shown in FIGS. 5 and 6, the outer circumferential surface of the end wall 120 may be formed so that the radial height of the connecting portion 123 connected to the leading edge 113 is the highest.

[0073] The connecting portion 123 portion A where the outer peripheral surface of the endwall 120 is connected to the radially inner end of the airfoil 110 may be formed so that the radial height on the leading edge 113 side is the highest, and the radial height of the connecting portion 123 portion B just before the trailing edge 114 is the lowest.

[0074] The radial height of the outer peripheral surface of the end wall 120 may be formed to be high from the first rim seal 121 to the portion connected to the leading edge 113, and then to be low toward the trailing edge 114.

[0075] As shown in Figures 5 and 6, the outer peripheral surface of the end wall 120 may include two recesses 127 near the corners of the end on the pressure face 111 side between the side of the leading edge 113 and the side of the trailing edge 114.

[0076] A convex portion 125-1 may be formed on the outer peripheral surface of the end wall 120 near a portion connected to the leading edge 113 on the pressure surface 111 side of the airfoil 110. In addition, a convex portion 125-2 may be formed on the outer peripheral surface of the end wall 120 at a portion slightly downstream from the middle portion on the pressure surface 111 side. The second convex portion 125-2 may be formed highest on the outer peripheral surface of the end wall 120 at a corner portion at an axial end portion higher than the lower end connecting portion 123 of the pressure surface 111.

[0077] The first recess 127-1 on the outer peripheral surface of the end wall 120 may be formed slightly downstream of the portion connected to the leading edge 113. The first recess 127-1 may be formed lowest at a corner portion of the axial end of the pressure surface 111, lower than the lower end connecting portion 123.

[0078] The second recess 127-2 may be formed on the outer peripheral surface of the end wall 120 slightly upstream of the connection with the trailing edge 114. The second recess 127-2 may be formed lowest in the corner portion of the axial end of the pressure surface 111, below the lower end connection portion 123.

[0079] The radial height of the outer peripheral surface of the end wall 120 may be formed to increase from the first rim seal 121 through the connecting portion 123 connecting the leading edge 113 to the first convex portion 125-1, then decrease to the second concave portion 127-2, and then increase again toward the rear end of the connecting portion 123 connecting the trailing edge 114.

[0080] As shown in FIG. 8, the outer peripheral surface of the end wall 120 may include two recesses 127 near the corners of the end on the suction surface 112 side between the side of the leading edge 113 and the side of the trailing edge 114.

[0081] When the corner portion of the axial end portion on the suction surface 112 side of the outer peripheral surface of the end wall 120 is used as a reference, the first convex portion 125-3, the first concave portion 127-3, the second convex portion 125-4, and the second concave portion 127-4 may be formed in order from upstream to downstream.

[0082] The first protrusion 125-3 may be located at the same position as the leading edge 113 or slightly upstream of the leading edge 113 with respect to the flow direction of the combustion gas.

[0083] The first recess 127-3 may be formed at the outer circumferential surface of the end wall 120 downstream from the leading edge 113, at the lowest point at the corner of the axial end at the junction with the suction surface 112.

[0084] The second protrusion 125-4 may be formed at a position slightly downstream from the middle portion on the suction surface 112 side. The second protrusion 125-4 may be formed highest at a corner portion of the outer peripheral surface of the end wall 120 that is closer to the axial end than the lower end connecting portion of the suction surface 112.

[0085] The second recess 127-4 may be formed lowest at the corner of the axial end at the junction with the trailing edge 114 and disposed between the junction with the second protrusion 125-4 and the trailing edge 114.

[0086] FIG. 9 is a top view showing the height of the outer peripheral surface of the end wall of a turbine blade according to one embodiment of the present invention, displayed with contour lines.

[0087] The outer peripheral surface of the endwall 120 of the turbine blade may have a greater radial height at the leading edge of the airfoil 110 and upstream of the pressure face. The outer peripheral surface of the endwall 120 may have a smaller radial height at the periphery where it meets the trailing edge of the airfoil 110.

[0088] The axial chord length of the turbine blade airfoil 110 is C x When this is done, the height of the outer peripheral surface of the end wall 120 in the radial direction is ±0.2C. x For example, C x When the height difference is about 100 mm, the outer peripheral surface of the end wall 120 can be formed in the range of -20 mm to +20 mm relative to the reference plane, and in particular, may be formed in the range of -10 mm to +15 mm. In Fig. 9, the contour lines can be displayed at intervals of 2 mm of height difference in the radial direction.

[0089] FIG. 10A is a photograph showing secondary vortices generated in a turbine blade according to the prior art, and FIG. 10B is a photograph showing secondary vortices generated in a turbine blade according to one embodiment of the present invention.

[0090] As shown in Figure 10A, in the turbine blade 10 according to the prior art of Figure 4A, a large secondary vortex is generated downstream of the leading edge 113 on the outer peripheral surface of the endwall 12 near the connection with the pressure surface 111.

[0091] In contrast, as shown in Figure 10B, the turbine blade 100 of the present invention shown in Figure 5 has significantly reduced secondary vortexes that occur immediately after the leading edge 113 on the outer surface of the end wall 120 near the connection with the pressure surface 111.

[0092] FIG. 11 is a graph showing pressure loss as a function of span position for a turbine blade according to the prior art and a turbine blade according to an embodiment of the present invention.

[0093] 11 shows the overall pressure loss coefficient at the outlet surface of the first stage turbine blade of the turbine. In the case of the turbine blade of the present invention, it can be seen that the overall pressure loss coefficient is reduced compared to the prior art, particularly in the span range of 0.1 to 0.3.

[0094] It has been confirmed that the turbine blade according to the present invention improves the stage efficiency of the first stage compared to the prior art, and also improves the overall efficiency of the turbine.

[0095] Furthermore, when the end wall shape of the present invention is applied, it has been confirmed that the efficiency of the gas turbine is improved, the efficiency of the combined cycle of the gas turbine and the steam turbine is improved by approximately 0.05%, the output of the gas turbine is improved, and the output of the combined cycle is improved, compared to the conventional technology.

[0096] According to the turbine blade of the present invention and a gas turbine including the same, by forming an asymmetrical endwall contour having multiple protrusions and multiple recesses from the rim seal on the leading edge side, secondary vortices can be reduced, thereby improving aerodynamic performance and efficiency.

[0097] Although one embodiment of the present invention has been described above, a person having ordinary skill in the art may modify and change the present invention in various ways by adding, changing, deleting or adding components within the scope of the concept of the present invention as set forth in the claims, and this also falls within the scope of the present invention.

Claims

1. an airfoil including a pressure surface, a suction surface, a leading edge, and a trailing edge; an endwall integrally formed on a lower portion of the airfoil; a root portion integrally formed on a lower portion of the end wall, The outer peripheral surface of the end wall is formed as a curved surface that is bent from a rim seal on one side to the other end of the turbine blade.

2. The end wall is a first rim seal extending upstream from the leading edge; The turbine blade according to claim 1 , further comprising: a second rim seal extending downstream from the trailing edge.

3. The turbine blade according to claim 2 , wherein the outer peripheral surface of the end wall is formed into a streamlined curved surface that connects from the first rim seal to the end portion on the trailing edge side.

4. The turbine blade according to claim 3 , wherein the second rim seal is positioned lower than the height of the outer peripheral surface of the endwall connected to the trailing edge.

5. The turbine blade according to claim 3 or 4, wherein the outer peripheral surface of the end wall is formed so that the height in the radial direction of the portion connected to the leading edge is the highest.

6. 6. The turbine blade according to claim 5, wherein the radial height of the outer peripheral surface of the end wall is increased from the first rim seal to a portion connected to the leading edge, and then decreased toward the trailing edge.

7. 7. The turbine blade of claim 6, wherein the outer peripheral surface of the endwall includes two recesses near corners of the pressure side end between the sides of the leading edge and the trailing edge.

8. 7. The turbine blade of claim 6, wherein the outer peripheral surface of the endwall includes two recesses near corners of the suction surface end between the sides of the leading edge and the trailing edge.

9. a compressor that draws in and compresses external air; a combustor that mixes fuel with the air compressed by the compressor and burns the fuel; a turbine in which turbine blades and turbine vanes are mounted inside a turbine casing, and the turbine blades are rotated by combustion gas discharged from the combustor, The turbine blade is an airfoil including a pressure surface, a suction surface, a leading edge, and a trailing edge; an endwall integrally formed on a lower portion of the airfoil; a root portion integrally formed on a lower portion of the end wall, A gas turbine in which the outer peripheral surface of the end wall is formed as a curved surface that is bent from a rim seal on one side to the other end.

10. The end wall is a first rim seal extending upstream from the leading edge; The gas turbine according to claim 9 , further comprising: a second rim seal extending downstream from the trailing edge.

11. The gas turbine according to claim 10 , wherein the outer peripheral surface of the end wall is formed into a streamlined curved surface that connects from the first rim seal to the end portion on the trailing edge side.

12. The gas turbine according to claim 11 , wherein the second rim seal is positioned lower than the height of the outer circumferential surface of the endwall connected to the trailing edge.

13. The gas turbine according to claim 11 or 12, wherein the outer peripheral surface of the end wall is formed so that a portion connected to the leading edge has the highest radial height.

14. 14. The gas turbine according to claim 13, wherein a radial height of an outer peripheral surface of the end wall is increased from the first rim seal to a portion connected to the leading edge, and then decreased toward the trailing edge.

15. 15. The gas turbine of claim 14, wherein the outer peripheral surface of the endwall includes two recesses near corners of the pressure side end between the sides of the leading edge and the trailing edge.

16. 15. The gas turbine of claim 14, wherein the outer peripheral surface of the endwall includes two recesses near corners of the suction surface end between the sides of the leading edge and the trailing edge.

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