JP6577400B2 - Turbine blade - Google Patents

Description

  The present invention relates to a turbine rotor blade, and more particularly, to a turbine rotor blade suitable for a plurality of inner peripheral side end wall portions to which airfoil portions are attached are arranged with a predetermined gap.

  In recent years, the number of blades of a turbine may be reduced from the viewpoint of cost reduction. However, when the number of blades is reduced, the load per blade tends to increase. In a blade with a large load on the blade, the flow in a cross section perpendicular to the flow of the mainstream gas in the vicinity of the end wall, regardless of the turbine central shaft side (inner periphery side) and the turbine casing side (outer periphery side), that is, Secondary flow increases. As the secondary flow increases, the flow rate in the vicinity of the end wall decreases, and the flow rate in the vicinity of the radial position on the outer peripheral side and inner peripheral side of the blade, that is, in the vicinity of the average diameter increases, and the blade load increases. As a result, it is known that the total pressure loss increases.

  In particular, in a turbine, cooling air may be guided from a compressor in order to cool a rotor portion to which turbine blades are attached. However, this cooling air is finally guided to a combustion gas passage so that combustion occurs. The gas flow will be disturbed. At this time, it is known that the cooling air mixed from the upstream side of the blade row further strengthens the secondary flow inside the blade row and further increases the total pressure loss.

  Also, in order to prevent an increase in total pressure loss in turbine blades with a heavy load on the turbine blades, the shape and arrangement of the concave area called the squealer on the outer peripheral side surface is devised, and the total pressure loss increases due to the flow through the gap. A method of reducing the total pressure loss by setting a tip shape called a winglet or a tip shape called a winglet on the outer peripheral side end face of the turbine rotor blade has been proposed (see Patent Document 1).

  Further, when attention is paid to the rotating shaft of the turbine rotor blade, Patent Document 2 discloses a turbine rotor blade in which a convex region is set in the end wall so as to be rotationally symmetric, that is, rotationally symmetric.

JP 2003-106104 A JP-A-11-148303

  However, in Patent Documents 1 and 2 described above, there is a concern that the performance of the turbine rotor blade is greatly affected by the mixed cooling air flow rate, and the application range is greatly affected by the blade height.

  Also, when the turbine blades are installed in a gas turbine, the disturbance of the cooling air in the flow field on the side farther from the turbine rotor (chip side) than the blades is the case of high-temperature gas. Has a great effect on the wings. In other words, the disturbance of the cooling air in the flow field increases the heat flow rate from the fluid side to the blades and increases the heat load, which may cause damage to the turbine blades.

  The present invention has been made in view of the above-described points, and the object of the present invention is to reduce the total pressure loss in the blade cross section on the tip side of the turbine rotor blade mixed with the cooling air, and suppress the performance deterioration due to the mixing of the cooling air. It is an object of the present invention to provide a turbine blade that can perform the above-described operation.

In order to achieve the above object, a turbine rotor blade of the present invention is attached to a rotary shaft, and is attached to an inner peripheral end wall portion located on the inner peripheral side of the rotary shaft, and to the inner peripheral end wall portion. An airfoil portion extending in a radial direction from the inner peripheral side end wall portion, and a plurality of the inner peripheral side end wall portions to which the airfoil portion is attached are arranged with a predetermined gap. A pressure surface having a concave curved surface shape in the chord direction, a suction surface having a curved surface shape convex in the chord direction, a blade leading edge and a blade trailing edge, and the blade thickness is on the blade leading edge side A turbine rotor blade formed so as to gradually increase toward the center side and gradually decrease from the middle toward the blade trailing edge side, and the pressure on the outer peripheral side of the airfoil portion A through-hole through the surface to the suction surface Is formed, the through hole is characterized in that the radial position towards the downstream side to the upstream side with respect to the through hole is formed lower.

  ADVANTAGE OF THE INVENTION According to this invention, the total pressure loss in the blade | wing cross section by the side of the tip of the turbine rotor blade with which cooling air mixes can be reduced, and the performance degradation by cooling air mixing can be suppressed.

1 is a partial cross-sectional view of a schematic configuration of a gas turbine to which a turbine rotor blade of the present invention is applied. It is a figure which shows the detail of arrangement configuration of the turbine rotor blade and turbine stationary blade of the gas turbine of this invention. It is a figure which expands and shows one turbine rotor blade of FIG. It is sectional drawing along the ZZ 'line of the airfoil part of FIG. It is a figure which shows the blade surface Mach number in the airfoil part surface cross section of the outer peripheral side blade end vicinity of the turbine rotor blade of this invention. It is a figure which shows blade surface Mach number distribution when the blade | wing load in the airfoil part surface cross section of the outer peripheral side blade edge part vicinity of the turbine rotor blade of this invention becomes small. It is a figure which shows the example of a turbine rotor blade comprised by the two rotor blades arrange | positioned adjacent to the turbine circumferential direction. It is the schematic which looked at the turbine rotor blade which concerns on Example 1 of this invention from the suction surface side. It is the schematic which looked at the turbine rotor blade which concerns on Example 1 of this invention from the pressure surface side. FIG. 10 is a cross-sectional view of a blade section cut by a surface 32 in FIGS. 8 and 9 and the cut surface viewed from the outside in the turbine radial direction. It is sectional drawing which follows the PQ line of FIG. It is a figure which shows the total pressure loss distribution of the turbine rotor blade which concerns on the cross section of the blade part in each position of a turbine radial direction.

  The turbine rotor blade of the present invention will be described below based on the illustrated embodiment. In each figure, the same symbols are used for the same components.

  FIG. 1 shows a schematic configuration of a gas turbine to which a turbine rotor blade of the present invention is applied.

  As shown in the figure, the gas turbine includes a turbine rotor 1 and a stator 2, and the turbine rotor 1 is mainly installed on the rotating shaft 3, and the turbine blade 4 and the compressor that are installed on the rotating shaft 3 and rotate together with the rotating shaft 3. The stator 2 includes a casing 7 and a turbine stationary blade 8 supported by the casing 7 and disposed so as to face the turbine rotor blade 4. It is configured.

  Compressed air and fuel from the compressor 5 are supplied to the combustor 6, and the fuel burns in the combustor 6 to generate high-temperature gas, and the generated high-temperature gas passes through the turbine stationary blade 8 and is a turbine blade. 4, and the turbine rotor 1 is driven via the turbine rotor blade 4.

  The turbine rotor blade 4 and the turbine stationary blade 8 that are exposed to the high-temperature gas are cooled as necessary. A part of the compressed air of the compressor 5 is used as the cooling medium for cooling.

  FIG. 2 shows details of the arrangement configuration of the turbine rotor blade 4 and the turbine stationary blade 8 of the gas turbine, and FIG. 3 shows one turbine rotor blade 4 in an enlarged manner.

  As shown in the figure, the turbine rotor blade 4 is attached to the turbine rotor 1 (see FIG. 1), and is located on the inner peripheral side with respect to the rotating shaft 3 of the turbine rotor blade 4, that is, on the inner periphery of the turbine rotor 1 side. The side end wall portion 10 and an airfoil portion 12 extending from the inner surface of the inner peripheral side end wall portion 10 in a direction in which the radial position increases (upward from the bottom of FIG. 3). Further, a gap 17 through which a fluid flows is formed between the closed surface 12c where the radius of the airfoil portion 12 is the largest and the outer surface of the outer peripheral end wall portion 16 that forms the gas flow path surface.

  The airfoil portion 12 has hollow portions 9a and 9b (described later), and a blade formed so as to cool the turbine rotor blade 4 from the inside by flowing a cooling medium through the hollow portions 9a and 9b. The mold part 12 may be configured. In FIG. 3, reference numeral 9 denotes an inlet of the cooling medium, and the cooling medium flows in the direction of the arrow to cool the airfoil part 12.

  As described above, the turbine rotor blade 4 is installed in the turbine rotor 1. However, as the cooling air supply source for cooling the turbine rotor blade 4, the compressor 5 is often used, and the cooling air is used as the turbine rotor 1. It is introduced into the turbine rotor blade 4 by using the cooling air introduction hole 14 provided in. The cooled cooling air is discharged from the discharge port 15 provided on the inner peripheral wall of the turbine rotor blade 4, and is eventually discharged to the gas path 20.

  Further, the air that has cooled the turbine rotor 1 is mixed into the gas flow path as indicated by arrows 18 and 19 from the gap between the turbine stationary blades 8A and 8B and the turbine rotor blade 4 on the inner peripheral side. However, since the turbine rotor blade 4 is rotating, it is known that especially the cooling air mixed by the arrow 18 from the upstream side strengthens the vortex developed from the inner peripheral side leading edge of the airfoil 12. This vortex effect dissipates energy between the wings.

  In FIG. 3, the line arrows indicate the flow of the cooling air, and the framed arrows indicate the flow of the hot gas, that is, the mainstream gas 11.

  FIG. 4 shows a cross-sectional shape of the above-described airfoil portion 12 along the line ZZ ′ in FIG.

  As shown in the figure, the airfoil portion 12 includes a pressure surface (blade side portion of the airfoil portion 12) 10b having a concave curved surface in the direction of the chord and a curved surface having a convex shape in the direction of the chord. And a blade leading edge 12a and a blade trailing edge 12b, and the blade thickness gradually increases from the blade leading edge 12a side toward the center side. And the blade thickness is gradually reduced from the middle toward the blade trailing edge 12b side. Further, the airfoil portion 12 has the above-described hollow portions 9a and 9b which are air cooling chambers, and a cooling medium is passed through the hollow portions 9a and 9b to cool the blade from the inside. .

  Further, fins (not shown) are provided in the hollow portions 9a and 9b at the front portion of the airfoil portion 12 in order to improve heat conversion. Then, similarly to the cooled turbine stationary blades 8A and 8B (see FIG. 3), the cooling air is discharged from the discharge port 15 provided in the inner peripheral wall, and is eventually discharged to the gas path 20.

  Such a cooling structure may be convection cooling or other cooling means. What is important is the shape of the outer peripheral blade end of the turbine rotor blade 4 mixed with cooling air.

  FIG. 5 is a diagram showing the blade surface Mach number in the airfoil surface cross section in the vicinity of the outer peripheral blade tip of the turbine blade 4 configured as described above. This figure shows the blade surface Mach number from the blade leading edge 12a to the blade trailing edge 12b of the suction surface 10a in the vicinity of the outer peripheral side end wall portion 16 by Ms, and the blade front of the pressure surface 10b in the inner peripheral side end wall portion 10 is shown. The blade surface Mach number from the edge 12a to the blade trailing edge 12b is indicated by Mp.

  As shown in FIG. 5, the blade surface Mach number of the suction surface 10a shows the maximum blade surface Mach number M_max at the intermediate portion between the blade leading edge 12a and the blade trailing edge 12b, and greatly decreases from the intermediate portion to the blade trailing edge 12b. ing. This is because the mainstream gas 11 is expanded when the mainstream gas 11 flows from the inlet to the outlet of the blade row constituted by the plurality of turbine rotor blades 4. M_min represents the minimum blade surface Mach number in the pressure surface 10b. As the difference between the maximum blade surface Mach number M_max and the minimum blade surface Mach number M_min increases, the difference between the maximum pressure and the minimum pressure acting on the airfoil portion 12 increases, and the load applied to the airfoil portion 12 increases. .

  In such a turbine rotor blade 4 with a large load on the airfoil portion 12, the flow in a cross section perpendicular to the flow of the mainstream gas 11 near the end wall regardless of the inner peripheral side or the casing 7 side, that is, Secondary flow increases. This is because in the vicinity of the end wall, the flow velocity decreases due to the influence of the viscosity of the mainstream gas 11, the centrifugal force acting on the mainstream gas 11 becomes smaller, and the pressure gradient becomes excessive. This is because the surface is attracted to the smaller suction surface 10a side.

  As the secondary flow increases, the flow rate of the mainstream gas 11 in the vicinity of the end wall decreases, and accordingly, the flow rate of the mainstream gas 11 in the vicinity of the average diameter increases and the blade load increases. An increase in blade load means an increase in the Mach number, and a friction loss on the wall surface, and a shock wave loss due to a shock wave in the supersonic region increases. As a result, the total pressure loss increases.

  FIG. 6 shows the blade surface Mach number distribution when the blade load becomes small. As shown in the figure, the difference between the maximum blade surface Mach number M′_max and the minimum blade surface Mach number M′_min when the blade load is reduced is the difference between the maximum blade surface Mach number_Mmax and the minimum blade surface Mach number M_min. It can be seen that it is smaller than the difference.

  In order to prevent such an increase in total pressure loss, in the turbine rotor blade 4, a method of suppressing the mixing loss of the flow generated in the gap between the turbine blade tip and the casing 7, or the blade leading edge 12 a of the airfoil portion 12. A method for suppressing the loss due to the secondary flow generated from the air has been proposed.

  In the method of suppressing the mixing loss of the flow generated in the gap between the turbine blade tip of the turbine rotor blade 4 and the casing 7, a concave portion called a squealer is set at the turbine blade tip to suppress the flow passing through the gap, Reducing mixing loss is the mainstream, and the new technology is to reduce the blade load near the blade tip by setting the blades in parallel in the flow direction called winglet and sharing the blade load between the two blades. And reducing the mixing loss by suppressing the flow through the gap.

  Hereinafter, the secondary flow vortex generated in the vicinity of the blade leading edge 12a is suppressed, and the secondary flow in the region sandwiched between the one suction surface and the other pressure surface of the two blades is also suppressed. A first embodiment of the turbine rotor blade 4 having the effect described above will be described with reference to FIGS. 7, 8, and 9.

  FIG. 7 shows an example of a turbine rotor blade composed of two rotor blades arranged adjacent to each other in the circumferential direction of the turbine, and suppresses an increase in secondary flow vortex generated in the vicinity of the blade leading edge 12a. In addition, this is an example of a turbine blade that also suppresses the secondary flow in the region sandwiched between the negative pressure surface 10a 'on one side and the pressure surface 10b on the other side of the two blades.

  FIG. 8 is a view of the turbine blade 4 according to the first embodiment of the present invention as viewed from the suction surface 10a side, and FIG. 9 is a view of the turbine blade 4 as viewed from the pressure surface 10b side.

  8 and 9, an arrow 13 represents the direction of gas flow. The blade leading edge 12a side is the upstream side, and the blade trailing edge 12b is the downstream side. R is a coordinate axis representing the radial position. An inner peripheral end wall portion 10 is located on the inner peripheral side of the airfoil portion 12. On the outer peripheral surface of the end portion of the airfoil portion 12, there is a through-hole 21 that is removed from the pressure surface 10 b side on the negative pressure surface 10 a side of the airfoil portion 12.

  That is, in the present embodiment, a through hole 21 is formed on the outer peripheral side of the airfoil portion 12 of the turbine rotor blade 4 so as to pass from the pressure surface 10b side to the negative pressure surface 10a side. As will be described, the radial position is lower from the downstream side toward the upstream side with respect to the through-hole 21.

  In this embodiment, the outer peripheral side of the airfoil portion 12 means the side (tip side) far from the turbine rotor 1 with respect to the airfoil portion 12 when the turbine rotor blade 4 is incorporated in a gas turbine. The inner peripheral side means the turbine rotor 1 side (hub side). Further, the outside means the outer peripheral side, and the inner means the inner peripheral side.

  More specifically, in the turbine rotor blade 4 of this embodiment, the above-described through hole 21 is located in the vicinity of the blade trailing edge 12b. Specifically, the blade leading edge 12a of the airfoil portion 12 located on the outer peripheral surface of the end portion of the airfoil portion 12 is set to 0%, and the blade trailing edge 12b is set to 100% on the basis of the coordinates in the rotation axis direction of the turbine. Then, the through hole 21 is formed so that the position thereof is within a range of 65% or more and 100% or less. This is because attention has been paid to the fact that an increase in blade load in the vicinity of the end wall causes the generation of vortices. In the region near the blade leading edge 12a of the turbine rotor blade 4, a secondary flow vortex that is wound up on the blade surface is generated. The mainstream gas also forms a vortex so as to be drawn into the secondary flow vortex.

  As shown in FIGS. 8 and 9, the turbine rotor blade 4 of the present embodiment has a through-hole 21 that extends to the negative pressure surface 10 a side on the pressure surface 10 b side on the outer peripheral side surface of the airfoil portion 12, that is, The through-hole 21 is provided through the through-hole inlet 33 to the through-hole outlet 34.

  Specifically, the contact between the inner peripheral side end wall part 10 and the blade leading edge 12a of the airfoil part 12 is 0%, and the contact point between the inner peripheral side end wall part 10 and the blade trailing edge 12b of the airfoil part 12 is provided. When 100%, the opening of the through hole 21 on the pressure surface 10b side exists between 10% and 50%, and the opening of the through hole 21 on the negative pressure surface 10a side is between 65% and 100%. It is formed to exist.

  This region is a region where the pressure gradient from the pressure surface 10b side to the negative pressure surface 10a side between the blades is the largest and the flow velocity is rapidly reduced, and vortices are likely to occur. By using the through hole 21 of the present embodiment, a flow from the pressure surface 10b side toward the negative pressure surface 10a side is generated, and the upstream airfoil portion 12, the downstream airfoil portion 12, and the configuration of the gap between them. The blade load on the end wall side can be reduced. Thereby, the secondary flow loss near the end wall can be suppressed.

  That is, when the position of the through hole 21 that passes from the pressure surface 10b side to the negative pressure surface 10a side is smaller than 15%, the differential pressure between the through hole inlet 33 and the through hole outlet 34 of the through hole 21 is small. Cannot flow to the suction surface 10a side, and the blade load reduction effect cannot be obtained. In addition, when the position of the through hole 21 that extends from the pressure surface 10b side to the negative pressure surface 10a side is larger than 50%, the differential pressure between the through hole inlet 33 and the through hole outlet 34 of the through hole 21 increases, The flow velocity of the fluid passing therethrough increases, and the mixing loss with the main flow increases on the suction surface 10a side. As a result, the loss increases on the suction surface 10a side, leading to performance degradation in the entire blade row.

  Next, the blade height direction of the through hole 21 will be described.

  The blade height position of the through hole 21 has a cross section of the airfoil portion 12 constituted by a gap between the two airfoil portions 12 on the outer peripheral side end face of the airfoil portion 12, and the inner peripheral end. When the wall portion 10 is defined as a blade height of 0% and the blade height at the tip of the outer peripheral airfoil portion 12 is defined as 100%, the gap between the two airfoil portions 12 is defined by the outer airfoil portion 12. Present (included) between 95% tip blade height and 100% blade height.

  That is, when the gap between the two airfoil portions 12 is 95% or less at the tip height of the airfoil portion 12 on the outer peripheral side, the hole of the through hole 21 becomes larger, so that the suction surface 10a from the pressure surface 10b side. As a result, the amount of work to be recovered from the mainstream air by the turbine rotor blade 4 is reduced. When the gap between the two airfoil portions 12 is larger than 95% of the blade height at the tip of the outer airfoil portion 12, the positional relationship with the turbine stationary blade 8 on the upstream side, and the wall boundary layer due to mixing of cooling air Due to the influence of the increase, the area of the through hole inlet 33 and the through hole outlet 34 of the through hole 21 passing from the pressure surface 10b side to the negative pressure surface 10a side is reduced, so that the flow rate passing through the through hole 21 is reduced and the airfoil portion The load reduction effect of 12 is reduced. Therefore, the apex of the outwardly convex shape is 95% or more and 100% or less.

  Further, in this embodiment, the contact between the inner peripheral side end wall portion 10 and the blade leading edge 12a of the airfoil portion 12 is 0%, and the inner peripheral side end wall portion 10 and the blade trailing edge 12b of the airfoil portion 12 are When the contact point is 100%, the fluid is mixed within the range of 0% or less and from the inner peripheral end wall 10 side.

  Next, a cross section of the blade tip portion of the turbine rotor blade 4 of the present embodiment will be described with reference to FIGS. 10 and 11. 10 shows a cross section of the turbine rotor blade 4 taken along the reference numeral 32 in FIGS. 8 and 9 as seen from the outer peripheral side of the inner peripheral side end wall portion 10, and FIG. 11 shows the PQ of FIG. Each section along the line is shown.

  In this figure, 33 is a through-hole inlet located on the pressure surface 10b side, and 34 is a through-hole outlet located on the negative pressure surface 10a side. Reference numeral 35 denotes a flow in the vicinity of the end wall from the through hole inlet 33 toward the through hole outlet 34. 36 is an upstream part on the inner peripheral side end wall part 10 side, and 37 is a downstream part on the inner peripheral side end wall part 10 side.

  As described above, in this embodiment, as shown in FIG. 10, a through hole 21 is formed on the outer peripheral side of the airfoil portion 12 of the turbine rotor blade 4 so as to pass from the pressure surface 10 b side to the negative pressure surface 10 a side. As shown in FIG. 11, the through hole 21 is formed with a lower radial position from the downstream side (P side) to the upstream side (Q side) with respect to the through hole 21.

  Thereby, due to the flow 35 from the through hole inlet 33 of the through hole 21 toward the through hole outlet 34, the differential pressure between the negative pressure surface 10a and the pressure surface 10b of the upstream side portion 36 on the inner peripheral side end wall portion 10 side is reduced. The load on the upstream wing is reduced. Further, the flow 35 from the through hole inlet 33 to the through hole outlet 34 of the through hole 21 reduces the flow velocity at the through hole outlet 34 as compared to the case without the through hole 21, so the inner end wall 10 side The differential pressure between the pressure surface 10b and the negative pressure surface 10a of the downstream portion 37 also decreases. By this principle, the blade load on the inner peripheral side end wall portion 10 side of the blade portion is reduced.

  FIG. 12 is a diagram showing the total pressure loss of the turbine rotor blade related to the cross section of the airfoil portion 12 at each position in the turbine radial direction. In the drawing, the tip side indicates the outside in the turbine radial direction (that is, the outer peripheral side end wall portion 16 side), and the hub side indicates the inner side in the turbine radial direction (that is, the inner peripheral side end wall portion 10 side).

  The total pressure loss according to this embodiment provided with the through hole 21 of this embodiment shown in FIGS. 8, 9, and 10 is drawn with a dotted line, and the total pressure loss according to the conventional example is drawn with a solid line. Yes.

  As can be seen from the figure, in the region close to the hub side, the total pressure loss according to the present example and the conventional example is substantially the same, but in the region close to the chip side, the total pressure loss according to the present example is the same. It can be seen that is reduced from the total pressure loss according to the conventional example. That is, in this embodiment, the total pressure loss in the blade cross section close to the outer peripheral side end wall portion 16 is reduced, and a more uniform total pressure loss is achieved in the vertical direction of the turbine radial direction. This means that more uniform expansion work is achieved over the entire area of the airfoil 12. Therefore, according to the present embodiment, the turbine efficiency can be improved and the fuel efficiency of the gas turbine can be reduced.

As described above, the turbine rotor blade 4 of the present embodiment has the through hole outlet 34 in the vicinity of the negative pressure surface 10a on the end face side of the turbine rotor blade 4 on the casing 7 side, and the through hole that induces the gas flow in the vicinity of the pressure surface 10b. Each inlet 33 is formed. By making the turbine rotor blade 4 in such a shape, in the main flow direction indicated by the arrow 13, it is possible to smoothly change the speed change while suppressing rapid deceleration and acceleration of the flow. Wings 4 can be provided.

  In the gas turbine configured as described above, the main flow fluid that flows in toward the turbine rotor blade 4 flows in from the blade leading edge 12a, flows along the airfoil portion 12, and flows out from the blade trailing edge 12b. By adopting the end shape of the turbine rotor blade 4 of the present embodiment, the secondary flow on the end side of the turbine rotor blade 4 is suppressed, and the inside of the mainstream fluid flowing along the suction surface 10a of the airfoil portion 12 is suppressed. The deceleration of the flow in the vicinity of the end wall on the circumferential side is suppressed, and the decrease in the Mach number on the suction surface 10a of the airfoil portion 12 of the turbine rotor blade 4 is also reduced.

  As a result, the total pressure loss in the blade cross section of the suction surface 10a of the airfoil portion 12 of the turbine rotor blade 4 can be reduced. Even when the load applied to the turbine rotor blade 4 is high or when the mixed cooling flow rate changes, the mechanism for generating the secondary flow is the same. An increase in total pressure loss can be suppressed.

  As described above, according to the present embodiment, in the turbine blade 4 in which the load on the blade increases and the secondary flow increases, the flow velocity of the gap flow passing through the blade tip gap of the turbine blade 4 is suppressed. The mixing loss of the main flow and the secondary flow can be reduced. Further, even when the tip clearance becomes zero, the velocity gradient in the flow direction is reduced due to the blade load reduction effect on the end side of the turbine rotor blade 4, the secondary flow vortex development is suppressed, and the secondary flow loss is reduced. To do.

  Therefore, according to the present embodiment, it is possible to provide the turbine rotor blade 4 having an effect of suppressing both the main flow and secondary flow mixing loss generated at the tip of the turbine rotor blade 4 and the secondary flow loss generated in the outer casing. .

  In addition, this invention is not limited to an above-described Example, Various modifications are included. For example, the above-described embodiments have been described in detail for easy understanding of the present invention, and are not necessarily limited to those having all the configurations described. Further, a part of the configuration of one embodiment can be replaced with the configuration of another embodiment, and the configuration of another embodiment can be added to the configuration of one embodiment. Further, it is possible to add, delete, and replace other configurations for a part of the configuration of each embodiment.

  DESCRIPTION OF SYMBOLS 1 ... Turbine rotor, 2 ... Stator, 3 ... Rotating shaft, 4 ... Turbine blade, 5 ... Compressor, 6 ... Combustor, 7 ... Casing, 8, 8A, 8B ... Turbine stationary blade, 9 ... Inlet, 9a, 9b ... Hollow part, 9c ... Pressure surface, 9f ... Air cooling chamber, 10 ... Inner end wall portion, 10a, 10a '... Negative pressure surface, 10b ... Pressure surface, 11 ... Mainstream gas, 12 ... Airfoil part, 12a ... blade leading edge, 12b ... blade trailing edge, 12c ... closed phase, 13 ... arrow, 14 ... cooling air introduction hole, 15 ... discharge port, 16 ... outer peripheral side end wall part, 17 ... gap, 18, 19 ... cooling air 20 ... Gas path, 21 ... Through hole, 32 ... Cross section line of turbine blade, 33 ... Through hole inlet, 34 ... Through hole outlet, 35 ... Flow from the through hole inlet to the outlet, 36 ... Upstream part of wing end wall side, 37 ... wing pea The downstream portion of Lumpur side.

Claims (4)

An inner peripheral end wall portion that is attached to the rotary shaft and is located on the inner peripheral side of the rotary shaft, and an airfoil portion that is attached to the inner peripheral end wall portion and extends in a radial direction from the inner peripheral end wall portion A plurality of the inner end wall portions to which the airfoil portion is attached are arranged with a predetermined gap, and the airfoil portion includes a pressure surface and a blade that form a concave curved surface in the chord direction. It consists of a suction surface that forms a curved surface convex in the chord direction, a blade leading edge and a blade trailing edge, and the thickness of the blade gradually increases from the blade leading edge side toward the center side, and from the middle A turbine rotor blade formed so as to gradually become smaller toward the blade trailing edge side,
A through hole extending from the pressure surface to the suction surface is formed on the outer peripheral side of the airfoil portion ,
The turbine rotor blade according to claim 1, wherein the through hole is formed such that a radial position thereof is lower from the downstream side toward the upstream side with respect to the through hole . The turbine rotor blade according to claim 1 ,
The airfoil part has an airfoil section formed by the through hole between the two airfoil parts on the outer peripheral side end surface of the airfoil part, and the inner peripheral side end wall part has a blade height of 0%, When the blade height at the tip of the airfoil portion on the outer peripheral side is defined as 100%, the through-hole between the two airfoil portions has a blade height of 95% from the blade height at the tip of the airfoil portion on the outer peripheral side. Turbine rotor blade characterized by being included between 100%. In the turbine rotor blade according to claim 1 or 2 ,
When the contact between the inner peripheral side end wall portion and the blade leading edge of the airfoil portion is 0%, and the contact between the inner peripheral side end wall portion and the blade trailing edge of the airfoil portion is 100%. Further, the opening of the through hole on the pressure surface side is present between 10% and 50%, and the opening of the through hole on the suction surface side is present between 65% and 100%. Turbine blades. In the turbine rotor blade according to any one of claims 1 to 3 ,
When the contact between the inner peripheral side end wall portion and the blade leading edge of the airfoil portion is 0%, and the contact between the inner peripheral side end wall portion and the blade trailing edge of the airfoil portion is 100%. In addition, a fluid is mixed in a range of 0% or less and from the inner peripheral side end wall portion side.

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