Turbine blades having non-linearly symmetric endwall contours and gas turbines 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 profile and a gas turbine including the same.

Background Art

[0002] A turbine is a mechanical device that obtains rotational force by impulse or reaction force using the flow of a compressible fluid such as steam or gas, and includes a steam turbine using steam and a gas turbine using high-temperature combustion gas.

[0003] Among these, a gas turbine is mainly composed of a compressor, a combustor, and a turbine. The compressor is provided with an air inlet for introducing air, and a plurality of compressor vanes and compressor blades are alternately arranged in a 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 alternately arranged in a turbine housing. Also, a rotor is arranged so as to penetrate the central portions of the compressor, the combustor, the turbine, and the exhaust chamber.

[0006] Both ends of the rotor are rotatably supported by bearings. A plurality of disks are fixed to the rotor, and at the same time, respective blades are connected, and a drive shaft such as a generator is connected to the end on the exhaust chamber side.

[0007] Such a gas turbine does not have a reciprocating mechanism like a piston of a four-stroke engine, so there is no mutual friction part like a piston-cylinder, the consumption of lubricating oil is extremely small, the amplitude, which is a characteristic of a reciprocating machine, is significantly reduced, and it has the merit of enabling high-speed movement.

[0008] To briefly explain the operation of a gas turbine, compressed air from a compressor is mixed with fuel and burned to produce high-temperature combustion gases, which are then injected into the turbine. As the injected combustion gases pass through the turbine vanes and blades, they generate rotational force, which causes the rotor to rotate.

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

[0010] On the other hand, when turbine blades are assembled on-site by workers, multiple turbine blades must be assembled from the first stage to the final stage, resulting in significant time and post-assembly tolerances. Furthermore, the uneven spacing or gaps between adjacent turbine blades can lead to problems with secondary vortices as combustion gases move.

[0011] When combustion gases pass through turbine blades, secondary vortices are generated, resulting in aerodynamic losses on the intake or pressure surface.

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

[0013] [Patent Document 1] Republic of Korea Published Patent Gazette No. 10-2019-0046118 (Published May 7, 2019) [Overview of the project] [Problems that the invention aims to solve]

[0014] The present invention aims to provide a turbine blade and a gas turbine including the same, which can reduce secondary vortices and improve aerodynamic performance by forming a non-linearly symmetric end wall contour having a plurality of convex portions and a plurality of concave portions from the rim seal on the leading edge side. [Means for solving the problem]

[0015] To achieve the above objective, the turbine blade of the present invention includes an airfoil having a pressure surface, an intake surface, a leading edge, and a trailing edge; an end wall integrally formed at the lower part of the airfoil; and a root portion integrally formed at the lower part of the end wall, wherein the outer circumferential surface of the end wall is formed as a curved surface that is bent from one rim seal 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 circumferential surface of the end wall may be formed as a streamlined curved surface that connects from the first rim seal to the trailing edge end.

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

[0019] The outer circumferential surface of the end wall may be formed with the highest radial height at the portion connected to the leading edge.

[0020] The radial height of the outer surface of the end wall may be formed such that it increases from the first rim seal to the portion connected to the leading edge, and then decreases towards the trailing edge.

[0021] The outer peripheral surface of the end wall can include two recesses near the corner of the end portion on the pressure surface side between the side portions of the leading edge and the side portions of the trailing edge.

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

[0023] In a gas turbine including a compressor that sucks and compresses external air, a combustor that mixes fuel with the air compressed by the compressor and burns the mixture, and 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 includes an airfoil including a pressure surface, a suction surface, a leading edge, and a trailing edge, an end wall integrally formed at a lower portion of the airfoil, and a root portion integrally formed at a lower portion of the end wall, and the outer peripheral surface of the end wall is formed as a curved surface bent from one side rim seal to the other end.

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

[0025] The outer peripheral surface of the end wall may be formed as a streamline curved surface connected from the first rim seal to the end portion on the trailing edge side.

[0026] [[ID=​​​​​​​The radial height of the outer surface of the end wall may be formed such that it increases from the first rim seal to the portion connected to the leading edge, and then decreases towards the trailing edge.

[0029] The outer circumferential surface of the end wall may include two recesses near the corner of the end on the pressure side, between the side of the leading edge and the side of the trailing edge.

[0030] The outer circumferential surface of the end wall may include two recesses near the corner of the end on the suction 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 and gas turbine including the present invention described above, by forming a non-linearly symmetric end wall contour having a plurality of protrusions and a plurality of recesses from the rim seal on the leading edge side, secondary vortices can be reduced and aerodynamic performance can be improved. [Brief explanation of the drawing]

[0032] [Figure 1] This is a partially cut perspective view of a gas turbine according to one embodiment of the present invention. [Figure 2] This is a cross-sectional view showing the schematic structure of a gas turbine according to one embodiment of the present invention. [Figure 3] This is a partial cross-sectional view showing the internal structure of a gas turbine according to one embodiment of the present invention. [Figure 4A] This is a partial perspective view showing a conventional turbine blade. [Figure 4B] This is a partial perspective view showing a turbine blade relating to one embodiment of the present invention. [Figure 5] This is a perspective view showing a turbine blade according to one embodiment of the present invention. [Figure 6] Figure 5 is a perspective view of the turbine blades from the pressure side. [Figure 7] Figure 5 is a perspective view of the turbine blade from the leading edge side. [Figure 8] Figure 5 is a perspective view of the turbine blades from the intake side. [Figure 9] This is a top view showing the height of the outer circumferential surface of the end wall of a turbine blade according to one embodiment of the present invention, indicated by contour lines. [Figure 10A] This is a photograph showing secondary vortices generated by a conventional turbine blade. [Figure 10B] This is a photograph showing secondary vortices generated in a turbine blade according to one embodiment of the present invention. [Figure 11] This graph shows the pressure loss according to the span position in a conventional turbine blade and a turbine blade according to one embodiment of the present invention. [Modes for carrying out the invention]

[0033] While the present invention can have various embodiments through diverse transformations, specific embodiments will be illustrated and described in detail in the detailed description. However, it should be understood that this does not intend to limit the present invention to specific embodiments, but rather includes all transformations, equivalents, or substitutions that fall within the spirit and technical scope of the present invention.

[0034] The terms used in this invention are used solely to describe specific embodiments and are not intended to limit the invention. Singular expressions include plural expressions unless the context clearly indicates otherwise. In this invention, terms such as “includes” or “having” are intended to specify the existence of features, figures, stages, operations, components, parts, or combinations thereof described in the specification, and should be understood not to preemptively exclude the possibility of the existence or addition of one or more other features, figures, stages, operations, components, parts, or combinations thereof.

[0035] Preferred embodiments of the present invention will be described in detail below with reference to the attached drawings. Note that, in the attached drawings, identical components are represented by the same reference numerals whenever possible. Furthermore, detailed descriptions of known functions and configurations that may obscure the gist of the present invention will be omitted. For similar reasons, some components in the attached drawings are exaggerated, omitted, or shown schematically.

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

[0037] As shown in Figure 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 comprises a plurality of radially arranged blades 1110. The compressor 1100 rotates the blades 1110, and air is compressed and moved by the rotation of the blades 1110. The size and installation angle of the blades 1110 can be changed depending on the installation position. In one embodiment, the compressor 1100 can be directly or indirectly connected to the turbine 1300 and can receive a portion of the power generated by the turbine 1300 and use it to rotate the blades 1110.

[0038] The air compressed in the compressor 1100 moves to the combustor 1200. The combustor 1200 includes a plurality of combustion chambers 1210 arranged in an annular manner and a fuel nozzle module 1220.

[0039] As shown in Figure 2, a gas turbine 1000 according to one embodiment of the present invention is equipped with a housing 1010, and a diffuser 1400 is provided on the rear side of the housing 1010 to discharge the combustion gas that has passed through the turbine. A combustor 1200 is positioned in front of the diffuser 1400 to receive compressed air and burn it.

[0040] To explain using the direction of airflow as a reference, the compressor section 1100 is located upstream of the housing 1010, and the turbine section 1300 is positioned downstream. Between the compressor section 1100 and the turbine section 1300, a torque tube unit 1500 is positioned as a torque transmission member that transmits the rotational torque generated in the turbine section 1300 to the compressor section 1100.

[0041] The compressor section 1100 is equipped with a plurality (for example, 14) of compressor rotor discs 1120, and each of the compressor rotor discs 1120 is fastened by tie rods 1600 so as not to separate in the axial direction.

[0042] Specifically, each of the compressor rotor discs 1120 is aligned axially with respect to each other, with a tie rod 1600, which constitutes the rotation axis, passing through approximately the center of each disc. Here, adjacent compressor rotor discs 1120 are positioned so that their opposing surfaces are pressed together by the tie rod 1600, making relative rotation impossible.

[0043] Multiple blades 1110 are radially connected to the outer circumferential surface of the compressor rotor disk 1120. Each blade 1110 is fastened to the compressor rotor disk 1120 and is equipped with a dovetail portion 1112.

[0044] Between each of the rotor discs 1120, there are vanes (not shown) fixed to the housing. Unlike the rotor discs, the vanes are fixed so as not to rotate and serve to align the flow of compressed air that has passed through the blades of the compressor rotor discs, guiding the air to the blades of the rotor disc located downstream.

[0045] The fastening method for the dovetail portion 1112 can be tangential or axial. This can be selected according to the required structure of the commercial gas turbine and can have the commonly known dovetail or fir-tree shape. In some cases, the blades can be fastened to the rotor disk using other fastening devices other than those described above, such as fasteners such as keys or bolts.

[0046] The tie rod 1600 is positioned to penetrate the centers of the plurality of compressor rotor discs 1120 and turbine rotor discs 1320, and the tie rod 1600 may consist of one or more tie rods. One end of the tie rod 1600 is fastened into the compressor rotor disc located on the upstream side, and the other end of the tie rod 1600 is fastened by a fixing nut 1450.

[0047] The form of the tie rod 1600 can consist of various structures depending on the gas turbine, and is not necessarily limited to the form shown in Figure 2. That is, as shown in the figure, it may have a form in which one tie rod penetrates the center of the rotor disc, or it may have a form in which multiple tie rods are arranged around the circumference, or a combination of these is also possible.

[0048] Although not shown in the diagram, a gas turbine compressor may have vanes that act as guide vanes located after the diffuser to adjust the fluid flow angle of the fluid entering the combustor inlet to the design flow angle after increasing the fluid pressure; these are called deswirlers.

[0049] In the combustor 1200, the incoming compressed air is mixed with fuel and burned to create a high-energy, high-temperature, high-pressure combustion gas, and the temperature of the combustion gas is raised to the heat resistance limit that the combustor and turbine components can withstand during the isobaric combustion process.

[0050] The combustors constituting the combustion system of a gas turbine may be arranged in multiples within a housing formed in a cell shape, and may consist of a burner including fuel injection nozzles, a combustor liner that forms a combustion chamber, and a transition piece that connects the combustor to the turbine.

[0051] Specifically, the liner provides a combustion space in which fuel injected by a fuel nozzle is mixed with compressed air from a compressor and burned. Such a liner may include a flame tube that provides a combustion space in which the fuel mixed with air is burned, and a flow sleeve that surrounds the flame tube and forms an annular space. A fuel nozzle is coupled to the front end of the liner, and a spark plug is coupled to the side wall.

[0052] Meanwhile, a transition piece is connected to the rear end of the liner to allow combustion gases, which are burned by the spark plug, to be sent to the turbine. The outer wall of such a transition piece is cooled by compressed air supplied from the compressor to prevent damage from the high temperature of the combustion gases.

[0053] For this purpose, the transition piece is provided with cooling holes that allow air to be injected into it. The compressed air cools the main body inside through the holes before flowing towards the liner.

[0054] Cooling air, which has cooled the transition piece described above, flows through the annular space of the liner, and compressed air provided outside the flow sleeve as cooling air can collide with the outer wall of the liner through cooling holes provided in the flow sleeve.

[0055] Meanwhile, the high-temperature, high-pressure combustion gases discharged from the combustor are supplied to the turbine 1300 described above. The supplied high-temperature, high-pressure combustion gases expand and collide with the turbine blades, generating a reaction force and producing rotational torque. This rotational torque is transmitted to the compressor via the torque tube described above, and any power exceeding the power required to drive the compressor is used to drive a generator or the like.

[0056] The turbine 1300 is basically similar in structure to that of a compressor. That is, the turbine 1300 is also equipped with a plurality of turbine rotor discs 1320 similar to the compressor rotor disc of a compressor. Therefore, the turbine rotor discs 1320 also include a plurality of turbine blades 1340 arranged radially. The turbine blades 1340 can also be coupled to the turbine rotor discs 1320 by a method such as a dovetail. In addition, turbine vanes 1330 fixed to the housing are provided between the blades 1340 of the turbine rotor discs 1320 to guide the flow direction of the combustion gases that have passed through the blades.

[0057] As shown in Figure 3, the turbine vanes 1330 are fixedly mounted within the housing by vane carriers 1335, which are endwalls connected to the inner and outer ends of the turbine vanes 1330. In contrast, a ring segment 1345 is mounted opposite 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. That is, the gap between the ring segment 1345 and the outer end of the turbine blade 1340 forms the tip clearance.

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

[0059] The conventional turbine blade 10 shown in Figure 4A includes an airfoil 11 and an end wall 12 integrally formed on the lower part of the airfoil. The airfoil 11 includes a pressure surface, an intake surface, a leading edge, and a trailing edge. The outer circumferential surface of the end wall 12 is formed in a substantially planar shape. More precisely, the outer circumferential surface of the end wall 12 is formed in a curved shape having a constant radius of curvature around the rotation axis of the turbine 1300.

[0060] The turbine blade 100 according to one embodiment of the present invention shown in Figure 4B includes an airfoil 110 and an end wall 120 integrally formed at the lower part of the airfoil. The end wall 120 may have rim seals formed on the axial upstream and downstream sides, respectively. In the present invention, the outer circumferential surface of the end wall 120 may be connected from the upper surface of the upstream rim seal to form a smooth curved surface.

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

[0062] A turbine blade 100 according to one embodiment of the present invention includes an airfoil 110 having a pressure surface 111, an intake surface 112, a leading edge 113, and a trailing edge 114, an end wall 120 integrally formed at the lower part of the airfoil, and a root portion 130 integrally formed at the lower part of the end wall.

[0063] The airfoil 110 includes a pressure surface 111 formed as a recess on one side, an intake surface 112 formed as 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 part of the airfoil 110, i.e., the radially inner end. The outer circumferential surface of the end wall 120 may be formed as a curved surface that is bent from one rim seal to the other end.

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

[0066] Multiple turbine blades 100 are mounted on the circumference of the turbine rotor disc 1320. The first rim seal 121 of the end wall 120 may be formed to seal between it and the fixed upstream turbine vane 1330, and the second rim seal 122 may be formed to seal between it and the fixed downstream turbine vane 1330.

[0067] The outer circumferential surface of the end wall 120 may be formed as 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 Figure 5, the upper surface of the end wall 120 may be formed as a smooth, streamlined curved surface that extends continuously from 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 end wall 120 that connects to the trailing edge 114.

[0070] The second rim seal 122 may be formed extending radially inward and downstream from the corner of the downstream end of the outer circumferential surface of the end wall 120.

[0071] As shown in Figure 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 circumferential surface of the end wall 120.

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

[0073] The connecting portion 123 A, where the outer circumferential surface of the end wall 120 connects to the radially inner end of the airfoil 110, may be formed with the highest radial height on the leading edge 113 side and the lowest radial height at the connecting portion 123 B immediately before the trailing edge 114.

[0074] The radial height of the outer surface of the end wall 120 may be formed such that it increases from the first rim seal 121 to the portion connected to the leading edge 113, and then decreases towards the trailing edge 114.

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

[0076] The outer circumferential surface of the end wall 120 may have a protrusion 125-1 formed near the portion connected to the leading edge 113 on the pressure surface 111 side of the airfoil 110. Alternatively, a protrusion 125-2 may be formed on the outer circumferential surface of the end wall 120 slightly downstream from the middle portion on the pressure surface 111 side. The second protrusion 125-2 may be highest on the outer circumferential surface of the end wall 120 at the corner portion at the axial end, above the lower end connection portion 123 of the pressure surface 111.

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

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

[0079] The radial height of the outer surface of the end wall 120 may be formed such that it increases from the first rim seal 121 through the connecting portion 123 that connects to the leading edge 113 to the first protrusion 125-1, then generally decreases to the second recess 127-2, and then increases again towards the rear end of the connecting portion 123 with the trailing edge 114.

[0080] As shown in Figure 8, the outer circumferential surface of the end wall 120 may include two recesses 127 near the corner of the end on the intake 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 on the suction surface 112 side of the outer circumferential 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 positioned at the same location as the leading edge 113, or slightly upstream of the leading edge 113, with reference to the flow direction of the combustion gas.

[0083] The first recess 127-3 may be formed lowest on the outer circumferential surface of the end wall 120, downstream of the leading edge 113, at the corner of the axial end where it connects with the intake surface 112.

[0084] The second protrusion 125-4 may be formed slightly downstream from the middle portion on the intake surface 112 side. The second protrusion 125-4 may be highest on the outer circumferential surface of the end wall 120 at the corner portion at the axial end, above the lower end connecting portion of the intake surface 112.

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

[0086] Figure 9 is a top view showing the height of the outer circumferential surface of the end wall of a turbine blade according to one embodiment of the present invention, indicated by contour lines.

[0087] The outer circumferential surface of the end wall 120 in the turbine blade may be formed with a higher radial height in the upstream region of the leading edge and pressure surface of the airfoil 110. The outer circumferential surface of the end wall 120 may be formed with a lower radial height in the peripheral region where it meets the trailing edge of the airfoil 110.

[0088] The axial chord length of the turbine blade airfoil 110 is C x In this case, the radial height of the outer surface of the end wall 120 is ±0.2C. x It can be formed within the range of C. x When the distance is approximately 100 mm, the outer surface of the end wall 120 can be formed in a range of -20 mm to +20 mm relative to the reference plane, and in particular, it may be formed in a range of -10 mm to +15 mm. In Figure 9, contour lines can be displayed at intervals of 2 mm for radial height differences.

[0089] Figure 10A is a photograph showing secondary vortices generated in a conventional turbine blade, and Figure 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, it can be seen that in the conventional turbine blade 10 shown in Figure 4A, a large secondary vortex is generated near the connection with the pressure surface 111 on the outer circumferential surface of the end wall 12, and downstream of the leading edge 113.

[0091] In contrast, as shown in Figure 10B, the turbine blade 100 according to the present invention in Figure 5 shows a significant reduction in secondary vortices generated immediately after the leading edge 113, near the connection point with the pressure surface 111 on the outer circumferential surface of the end wall 120.

[0092] Figure 11 is a graph showing the pressure loss according to the span position in a conventional turbine blade and a turbine blade according to one embodiment of the present invention.

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

[0094] We have confirmed that the turbine blade according to the present invention improves the stage efficiency of the first stage compared to the conventional technology, and also improves the overall efficiency of the turbine.

[0095] Furthermore, when the endwall shape of the present invention is applied, it was confirmed that the efficiency of the gas turbine is improved compared to the conventional technology, the efficiency of the combined cycle of the gas turbine and 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.

[0096] According to the turbine blade of the present invention and the gas turbine including it, by forming a non-linearly symmetric end wall contour having a plurality of protrusions and a plurality of recesses from the rim seal on the leading edge side, secondary vortices can be reduced and aerodynamic performance and efficiency can be improved.

[0097] Although one embodiment of the present invention has been described above, any person with ordinary skill in the art can modify and change the present invention in various ways, such as by adding, changing, deleting, or adding components, without departing from the spirit of the invention as described in the claims, and this is also included within the scope of the rights of the present invention.

Claims

1. An airfoil including a pressure surface, an intake surface, a leading edge, and a trailing edge, An end wall integrally formed at the lower part of the airfoil, The end wall includes a root portion integrally formed at the lower part of the end wall, The outer circumferential surface of the end wall is formed as a curved surface that is bent from one rim seal to the other end. The outer circumferential surface of the end wall includes two recesses located between the side of the leading edge and the side of the trailing edge, and near the corner of the end on the pressure surface side. The two recesses include a first recess on the upstream side and a second recess on the downstream side. The radial height of the outer surface of the end wall is lower than the height of the first recess at the location of the second recess. Turbine blades.

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

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

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

5. The turbine blade according to claim 3 or 4, wherein the outer circumferential surface of the end wall is formed to have the highest radial height at the portion connected to the leading edge.

6. The turbine blade according to claim 5, wherein the radial height of the outer circumferential surface of the end wall is formed to increase from the first rim seal to the portion connected to the leading edge, and then decrease towards the trailing edge.

7. The turbine blade according to claim 6, wherein the outer circumferential surface of the end wall includes two recesses near the corner of the end on the intake side, between the side of the leading edge and the side of the trailing edge.

8. The turbine blade according to claim 1, wherein the outer circumferential surface of the end wall includes a first protrusion located upstream of the first recess on the pressure surface side and a second protrusion located between the first recess and the second recess, and the radial height of the outer circumferential surface of the end wall is lower than the first protrusion at the position of the first recess and lower than the second protrusion at the position of the second recess.

9. A compressor that takes in outside air and compresses it, A combustor that mixes fuel with the air compressed by the aforementioned compressor and burns it, In a gas turbine, which includes a turbine in which turbine blades and turbine vanes are mounted inside a turbine casing, and the turbine blades are rotated by combustion gases discharged from the combustor, The turbine blades are An airfoil including a pressure surface, an intake surface, a leading edge, and a trailing edge, An end wall integrally formed at the lower part of the airfoil, The end wall includes a root portion integrally formed at the lower part of the end wall, The outer circumferential surface of the end wall is formed as a curved surface that is bent from one rim seal to the other end. The outer circumferential surface of the end wall includes two recesses located between the side of the leading edge and the side of the trailing edge, and near the corner of the end on the pressure surface side. The two recesses include a first recess on the upstream side and a second recess on the downstream side. The radial height of the outer surface of the end wall is lower than the height of the first recess at the location of the second recess. Gas turbine.

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

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

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 end wall connected to the trailing edge.

13. The gas turbine according to claim 11 or 12, wherein the outer circumferential surface of the end wall is formed to have the highest radial height at the portion connected to the leading edge.

14. The gas turbine according to claim 13, wherein the radial height of the outer surface of the end wall is formed to increase from the first rim seal to the portion connected to the leading edge, and then decrease towards the trailing edge.

15. The gas turbine according to claim 14, wherein the outer circumferential surface of the end wall includes two recesses between the side of the leading edge and the side of the trailing edge, near the corner of the end on the intake side.

16. The turbine blade according to claim 1, wherein the outer circumferential surface of the end wall includes a first protrusion located upstream of the first recess on the pressure surface side and a second protrusion located between the first recess and the second recess, and the radial height of the outer circumferential surface of the end wall is lower than the first protrusion at the position of the first recess and lower than the second protrusion at the position of the second recess.

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