JP5010507B2 - Turbine stage of axial flow turbomachine and gas turbine - Google Patents

Description

  The present invention relates to a turbine stage used in an axial turbomachine such as a gas turbine.

  In an axial-flow turbomachine such as a gas turbine or a jet engine, a moving blade is usually arranged on the outer peripheral side of the rotor, and a stationary blade is arranged on the inner peripheral side of the shroud so as to face the moving blade. There are two turbine stages. In such a turbine stage for an axial-flow turbomachine, a member to which blade ends are joined, that is, a so-called platform outer wall surface (hereinafter referred to as an end wall) is provided with irregularities so that adjacent blades There has been proposed a technique for improving the aerodynamic performance such as pressure loss by improving the gas flow in the flow path (see, for example, Patent Document 1).

  In Patent Document 1 below, a protruding portion (bulge) that protrudes outward (flow channel side) from the reference surface is formed in a portion of the end wall that is close to the leading edge of the wing's abdominal surface (concave pressure side). In addition, a turbine stage has been proposed in which a recess recessed inward from the reference surface is formed in a portion close to the rear edge of the rear side of the blade (convex suction side).

US Pat. No. 7,134,842

  By the way, in the turbine stage as in Patent Document 1, attention is paid only to the aerodynamic performance in the flow path (hereinafter simply referred to as “flow path between blades”) in which the shape of the end wall is formed between adjacent blades. Therefore, a part having a high heat transfer rate may be locally generated in the end wall.

  If the area of the part with a high heat transfer coefficient is large in this way, the amount of heat transferred from the gas flowing through the flow path between the blades to the platform and the wing through the part increases, and the thermal load received by the platform and the wing increases. Problem arises. When the thermal load increases, it becomes necessary to increase the flow rate of the refrigerant that cools the platform and blades, which may reduce the efficiency of the turbine. It becomes difficult to apply to the stage.

  Therefore, in the turbine stage, there is a demand for an end wall shape that can suppress the occurrence of a portion having a high heat transfer coefficient in the end wall while improving the aerodynamic performance in the inter-blade flow path.

  The present invention has been made in view of the above, and an object of the present invention is to provide a turbine stage and a gas turbine capable of reducing a thermal load applied to a platform and blades.

In order to achieve the above object, a turbine stage according to the present invention is used in an axial-flow turbomachine, and a blade extending in a radial direction of a rotating shaft from a platform extending in a circumferential direction of the rotating shaft has a rotating shaft. Are arranged in the circumferential direction of the turbine, and an inter-blade flow path is formed between adjacent blades. The end wall, which is the wall surface of the platform on the inter-blade flow path side, is recessed in the radial direction of the rotating shaft. A recessed portion is provided, and the recessed portion extends along the front edge side of the back surface of the other wing facing the abdominal surface and the abdominal side recessed portion extending along the front edge side of the abdominal surface of one wing of adjacent wings. It has an extending dorsal recess, and a connecting recess for connecting the dorsal recess and the ventral recess to the downstream side of the inter-blade channel .

  The turbine stage which concerns on this invention WHEREIN: A recessed part shall have a front edge side recessed part extended along a back surface from the front edge vicinity separately from the said back side recessed part.

  In the turbine stage according to the present invention, the back-side recess and the ventral-side recess are recessed with an equal distance in the radial direction of the rotating shaft with respect to a reference surface extending in the circumferential direction of the rotating shaft of the end wall. Can be.

A gas turbine according to the present invention is a gas turbine including a compressor that takes in and compresses combustion air, a combustor that burns the compressed air, and a turbine that extracts mechanical power from the combustion gas. Of the turbine stages constituting the turbine, an end wall that is a wall surface on the inter-blade flow path side of the platform is provided with a concave portion that is recessed in the radial direction of the rotating shaft, and the concave portion is one of adjacent blades. A ventral recess extending along the front edge of the abdominal surface of the wing, a back recess extending along the front edge of the back of the other wing facing the abdominal surface, and a back recess and a vent recess And a connection recess connected to these on the downstream side of the flow path between the blades .

  In the turbine stage according to the present invention, the concave portion has a connection concave portion that connects the back side concave portion and the ventral side concave portion to the downstream side of the flow path between the blades. Combustion gas from the recesses can be merged to reduce the flow of combustion gas toward the throttle part of the inter-blade channel between the trailing edge of one blade and the back of the other blade. It can suppress that the area | region with a high heat transfer rate arises in the end wall which faces a part.

  In the turbine stage according to the present invention, the recess has a front edge side recess extending from the vicinity of the front edge along the back surface separately from the back side recess, so that the flow of combustion gas toward the front edge side of the back surface The high-speed combustion gas flow can be prevented from colliding with the front edge side of the rear surface and turbulence in the combustion gas flow.

Further, in the gas turbine according to the present invention, a recess recessed in the radial direction of the rotating shaft is provided on an end wall which is a wall surface on the inter-blade flow path side of the platform among the turbine stages constituting the turbine. A ventral recess extending along the front edge of the ventral surface of one of the adjacent wings, a back recess extending along the front edge of the back of the other wing facing the ventral surface, The side recess and the ventral recess are connected to the downstream side of the flow path between the blades , so that the combustion gas from the ventral recess and the back recess is merged, The flow of combustion gas toward the throttle part of the inter-blade channel between the trailing edge of the other blade and the back surface of the other blade can be reduced, and the heat transfer coefficient of the end wall facing the throttle part is reduced. Generation of a high region can be suppressed.

  Hereinafter, the present invention will be described in detail with reference to the drawings. Note that the present invention is not limited to the embodiments. In addition, constituent elements in the following embodiments include those that can be easily assumed by those skilled in the art or those that are substantially the same.

  A configuration of a turbine stage for an axial-flow turbomachine according to the present embodiment will be described with reference to FIGS. FIG. 1 is a development view showing the shape of an end wall in a turbine stage according to the present embodiment. FIG. 2 is a perspective view showing a configuration example of the turbine stage. FIG. 3A is a cross-sectional view taken along line AA in FIG. 3-2 is a cross-sectional view taken along line BB in FIG. 3C is a cross-sectional view taken along the line CC of FIG. FIG. 4 is a diagram showing a static pressure distribution in a flow path between blades of a turbine stage according to the related art. FIG. 5 is a diagram illustrating a static pressure distribution in the inter-blade flow path of the turbine stage according to the present embodiment. FIG. 6 is a diagram showing the distribution of heat transfer coefficient in the end wall of the turbine stage according to the prior art. FIG. 7 is a diagram showing the heat transfer coefficient in the end wall of the turbine stage according to this embodiment. FIG. 8 is a cross-sectional view showing the overall configuration of the gas turbine according to the present embodiment.

  1 and 4 to 7 are drawings in which a platform extending in the circumferential direction (indicated by an arrow C in the drawing) of the rotary shaft of the rotor of the axial-flow turbomachine is linearly developed. Moreover, in FIG. 1, FIG. 4-FIG. 7, about the wing | blade, only the airfoil (aerofoil) which is the cross section is shown.

  As shown in FIG. 8, the gas turbine 1 is provided with an air intake 2, a compressor 3, a combustor 4, and a turbine 5 from the upstream side to the downstream side of the air flow. The air taken in from the air intake 2 is compressed by the compressor 3 and is supplied to the combustor 4 as high-temperature and high-pressure compressed air. The combustor 4 supplies gas fuel such as natural gas or liquid fuel such as light oil and heavy oil to the compressed air in a combustion chamber formed therein, and burns the fuel to generate high-temperature and high-pressure combustion gas. Generate. High-temperature and high-pressure combustion gas generated in the combustion chamber is injected from the combustor 4 toward the turbine 5. The turbine 5 is rotationally driven by the injected combustion gas. Part of the mechanical power received by the turbine 5 from the combustion gas is used for rotational driving of the compressor 3, and the rest is used for power generation and the like.

  Such a turbine 5 usually has a casing 7 (chamber) in which stationary blades are arranged on the inner periphery, and a rotor 8 in which moving blades are arranged on the outer periphery, and the rotor 8 includes combustion gas. In response to this flow, it rotates together with the moving blade. In the turbine 5, the blade row of the moving blades and the blade row of the stationary blades constitute a turbine stage as a pair. That is, the turbine 5 is composed of a plurality of turbine stages. One of the moving blades and the stationary blades constituting the turbine stage is simply referred to as “blade”. That is, the turbine stage includes at least one of a blade row of moving blades and a blade row of stationary blades.

  As shown in FIG. 2, in the turbine stage 10, the platform 14, which is a member to which the blades 20 are coupled, is an axis of a rotor rotation axis (hereinafter simply referred to as “rotation axis”) as indicated by an arrow A in the figure. It has a width in the direction and extends in the circumferential direction of the rotating shaft (indicated by an arrow C in the figure). On the platform 14, wings of the same shape are arranged over the entire circumference with a predetermined interval in the circumferential direction of the rotation shaft. That is, the turbine stage 10 is provided with blade rows forming a row in the circumferential direction of the rotor rotation shaft.

  The outer surface of each blade 20 has a front surface 22 and a rear surface 28 as boundaries, and a front surface 24 that receives pressure from combustion gas (flow direction is indicated by an arrow G) flowing into the turbine stage 10, and a front surface 24. The back surface 26 is formed so as to face the surface. In other words, the wing 20 is surrounded by the abdominal surface 24 and the back surface 26 to define its wing shape. The end of each blade 20 is coupled to the platform 14 and arranged in the circumferential direction of the rotation axis.

  Thus, a flow path (hereinafter referred to as a flow path between blades) 30 through which combustion gas flows is formed between adjacent blades 20. In the following description, the wall surface of the platform 14 on the side of the inter-blade channel 30 is referred to as “end wall” and indicated by reference numeral 15. In the lower part of FIG. 2, the intersection line between the end wall 15 and the abdominal surface 24 and the back surface 26 of the wing 20, that is, the airfoil shape is indicated by a two-dot chain line.

  In the following description, the upstream side in the flow direction of the combustion gas in the axial direction of the rotating shaft indicated by the arrow A in the figure is referred to as “front”. On the other hand, the downstream side in the flow direction of the combustion gas is described as “rear”.

  In the turbine stage 10, the end wall 15 extends in the circumferential direction of the rotating shaft (indicated by an arrow C in the drawing). The end wall 15 has a recess (hereinafter referred to as a recess) that is recessed inward in the radial direction of the rotation axis with respect to a surface (hereinafter referred to as a reference plane) 15a located at a distance equal to the radial direction from the rotation axis. 50) is formed. In the figure, the radial direction of the rotating shaft is a direction orthogonal to both the axial direction indicated by arrow A and the circumferential direction indicated by arrow C. The recesses 50 are provided between adjacent blades 20 and are arranged over the entire circumference in the circumferential direction of the rotating shaft.

  As shown in FIGS. 1 and 2, an inter-blade channel 30 is formed between the abdominal surface 24 of one wing 20 and the back surface 26 of the other wing 20 facing the abdominal surface 24 of the wing 20. The recess 50 is provided corresponding to the inter-blade channel 30 in the end wall 15. The concave portion 50 has a substantially V-shape in which the front edge 22 side, that is, the front side is up, and the rear edge 28 side, that is, the rear side, is down in the axial direction A of the rotating shaft.

  In the inter-blade channel 30, the upstream side in the combustion gas flow direction, that is, the front edge 22 side is simply referred to as “upstream side”. On the other hand, the downstream side in the flow direction of the combustion gas, that is, the rear edge 28 side is simply referred to as “downstream side”. In the inter-blade channel 30, the narrowest part of the channel width formed between the trailing edge 28 of one blade 20 and the back surface 26 of the other blade 20 is hereinafter referred to as a “throttle portion”. This is indicated by reference numeral 33.

  Further, in FIG. 1, a portion of the recess 50 having the same distance (hereinafter referred to as “depth”) recessed in the radial direction of the rotary shaft with respect to the reference surface 15 a located at an equal distance in the radial direction from the rotary shaft. Are connected by the same line. Each line is given an arbitrary index (0 to 3) indicating the depth. The index 0 (zero) indicates the reference surface 15a, and indicates that the depth increases (deeply) as the index increases. The deepest part of the recess 50 is about 4 to 6 mm.

  As shown in FIGS. 1 and 3-1 to 3-3, the recess 50 includes a ventral recess 54 extending along the front edge 22 side of the ventral surface 24 of the one wing 20 and the other wing. The back side recess 56 extending along the front edge 22 side portion of the rear surface 26 of the back surface 20, and the abdominal side recess 54 and the back side recess 56 are separated from each other with respect to the abdominal side recess 54 and the back side recess 56. It has the connection recessed part 58 connected in the downstream of the flow path 30. FIG.

  As shown in FIGS. 1, 3-1, and 3-2, the ventral recess 54 extends along the ventral surface 24 of the wing 20 and does not cover the ventral surface 24. The recesses 54 are configured so that no irregularities are formed on the abdominal surface 24. The front end 54 a in the axial direction A of the ventral recess 54 is set slightly forward compared to the front edge 22. The abdominal recess 54 configured as described above can reduce the flow velocity of the combustion gas that flows toward the abdominal surface 24 of the blade 20 out of the combustion gas flowing into the inter-blade channel 30 along the end wall 15. ing.

  On the other hand, the dorsal recess 56 extends along the back surface 26 of the other wing 20 adjacent to the one wing 20, similarly to the ventral recess 54. The back-side recess 56 is configured not to cover the back surface 26, that is, the back-side recess 56 does not form unevenness on the back surface 26. The front end 56 a in the axial direction A of the back side recess 56 is set slightly forward compared to the front edge 22. The back-side concave portion 56 configured in this way can reduce the flow velocity of the combustion gas that flows toward the back surface 26 of the blade 20 out of the combustion gas flowing into the inter-blade channel 30 along the end wall 15. ing.

  As shown in FIGS. 3A and 3B, the back-side concave portion 56 and the ventral-side concave portion 54 are arranged in the radial direction of the rotation shaft (see FIG. Indented by a distance equal to the arrow R). That is, the bottom 56c of the back-side recess 56 and the bottom 54c of the abdominal-side recess 54 are set to be located at the same distance from the reference surface 15a in the rotational axis radial direction. A connection recess 58 is provided on the rear side in the axial direction A with respect to the abdominal recess 54 and the back recess 56, that is, on the downstream side of the inter-blade channel 30, that is, on the rear edge 28 side. The recess 56 is connected through the connection recess 58.

  As shown in FIGS. 1 and 3-3, the connection recess 58 extends along the inter-blade channel 30 formed between the ventral surface 24 of one wing 20 and the back surface 26 of the other wing 20. The rear end 58a of the connection recess 58 in the axial direction A is set in the vicinity of the trailing edge 28 of the wing 20. The connection recess 58 configured as described above joins the combustion gas from the abdominal recess 54 and the back recess 56 to form a space between the trailing edge 28 of one blade 20 and the back surface 26 of the other blade 20. It is possible to guide to the throttle part 33 of a certain inter-blade channel 30 and to reduce the flow of combustion gas from the front edge 22 side toward the throttle part 33 in the inter-blade channel 30. ing.

  Next, the static pressure distribution in the inter-blade flow path of the turbine stage according to the present embodiment will be described with reference to FIGS. 4 and 5 in comparison with a turbine stage in which no recess is formed. In FIG.4 and FIG.5, the part with equal static pressure is connected with the same line (pressure contour line). Each line is given an arbitrary index (2 to 11) indicating a static pressure, and indicates that the static pressure increases as the index increases. In addition, it is about 90 kPa in the site | part where static pressure is the lowest.

  In the turbine stage 100 that does not have the recess 50, as shown in FIG. 4, the static pressure gradually decreases from the front edge 22 side toward the throttle portion 33 in the inter-blade channel 30. In particular, on the upstream side of the throttle portion 33, as indicated by an index 8 to an index 4, the static pressure is reduced with a substantially uniform pressure gradient.

  On the other hand, in the turbine stage 10 in which the recess 50 is formed, as shown in FIG. 5, the pressure contour lines of index 8 to index 4 are denser than the turbine stage 100 not having the recess 50, and the throttle part The pressure gradient immediately upstream of 33 is steep. Accordingly, the pressure gradient is gentle on the upstream side (front edge 22 side) of the inter-blade channel 30, and the flow of combustion gas flowing into the inter-blade channel 30 collides with the abdominal surface 24 or the back surface 26. Thus, it is possible to suppress the occurrence of a turbulent flow (hereinafter referred to as a secondary flow).

Next, the distribution of the heat transfer coefficient in the end wall of the turbine stage according to the present embodiment will be described with reference to FIGS. 6 and 7 in comparison with the turbine stage in which no recess is formed. In FIG.6 and FIG.7, the part with the same heat transfer coefficient [W / (m < ² > * K)] is connected with the same line (isothermal transfer coefficient line). Each line is given an arbitrary index (1 to 5) indicating the heat transfer coefficient, indicating that the heat transfer coefficient increases as the index increases.

  In the turbine stage 100 having no recess, as shown in FIG. 6, a region having a particularly high heat transfer rate in the throttle portion 33 of the inter-blade channel 30 compared to the other, for example, an isothermal transfer indicated by an index of 5 A region surrounded by a rate line is generated. The portion extends along the back surface 26 from the throttle portion 33. When such a region having a high heat transfer coefficient is generated over a wide area in the end wall 15, the heat of the combustion gas flowing through the inter-blade channel 30 is transmitted from the portion to the platform 14, and is coupled to the platform 14. This causes a problem that the thermal load received by the blade 20 is increased.

  On the other hand, in the turbine stage 10 according to the present embodiment, since the recesses 50 are formed in the end wall 15, as shown in FIG. 7, the turbine without the recesses 50 in the throttle portion 33 of the inter-blade channel 30. Compared to the stage 100, the area of the high heat transfer coefficient region surrounded by the isothermal transfer coefficient line of index 5 is reduced. Further, in the inter-blade channel 30, the area surrounded by the isothermal transfer coefficient lines of index 3 and index 2 is reduced, and the area surrounded by the isothermal transfer coefficient line of index 1 is increased accordingly. ing.

  In this way, the turbine stage 10 in which the recess 50 is formed has a reduced area of high heat transfer rate in the end wall 15 corresponding to the inter-blade channel 30 as compared with the turbine stage 100 that does not have the recess 50. Yes. As a result, the turbine stage 10 can reduce the amount of heat transferred from the combustion gas flowing through the inter-blade passage 30 to the platform 14 via the end wall 15, and the platform 14 and the blades coupled thereto. Twenty thermal loads can be reduced.

  As described above, in the turbine stage 10 according to this embodiment, the end wall 15 that is the wall surface of the platform 14 on the side of the inter-blade flow path 30 has the concave portion 50 that is recessed in the radial direction R of the rotating shaft in the circumferential direction C. The recesses 50 are arranged such that the abdomen recesses 54 extending along the front edge 22 side of the abdomen surface 24 of one wing of the adjacent wings 20 and the other wings facing the abdomen surface 24. A back side recess 56 extending along the front edge 22 side of the back surface 26.

  In the turbine stage 10, out of the combustion gas flowing into the inter-blade channel 30 along the end wall 15, the back-side recess 56 decelerates the flow velocity of the combustion gas toward the rear surface 26 of the blade 20, and the ventral surface 24 of the blade 20. The abdominal recess 54 decelerates the flow rate of the combustion gas toward. Thus, the turbine stage 10 can reduce the occurrence of a region having a high heat transfer coefficient in the end wall 15 facing the inter-blade channel 30 as compared with the turbine stage 100 that does not have the recess 50. The thermal load that the blade 20 and the platform 14 receive from the combustion gas flowing through the flow path 30 can be reduced.

  Further, in the turbine stage 10 according to the present embodiment, the concave portion 50 has a connection concave portion 58 that connects the back side concave portion 56 and the ventral side concave portion 54 to the downstream side of the inter-blade channel 30 with respect to these. Therefore, the combustion gas from the abdominal recess 54 and the back recess 56 is merged, and the throttle portion of the inter-blade channel 30 between the trailing edge 28 of one blade 20 and the back surface 26 of the other blade 20. The flow of the combustion gas toward 33 can be decelerated. It can suppress that the area | region with a high heat transfer rate arises in the end wall 15 which faces the said throttle part 30. FIG.

  In the turbine stage 10 according to this embodiment, the back-side recess 56 and the abdomen-side recess 54 are arranged in the radial direction R of the rotating shaft with respect to the reference surface 15a extending in the circumferential direction C of the rotating shaft. Although it was assumed that it was dented with equal distance, it is not limited to this. The distance in which the back side concave portion 56 and the ventral side concave portion 54 are recessed in the radial direction of the rotating shaft can be set to an arbitrary value according to the aerodynamic performance or the like in the inter-blade channel.

  The gas turbine 1 according to this embodiment includes a compressor 3 that takes in and compresses combustion air, a combustor 4 that combusts the compressed air, and a turbine 5 that extracts mechanical power from the combustion gas. Among the turbine stages 10 constituting the turbine 5, the end wall 15, which is the wall surface of the platform 14 on the inter-blade channel 30 side, is provided with a recess 50 that is recessed in the radial direction of the rotating shaft. It has been. The concave portion 50 includes a ventral recess 54 extending along the front edge 22 side of the abdominal surface 24 of one of the adjacent wings 20, and the front edge 22 of the back surface 26 of the other wing facing the abdominal surface 24. And a back recess 56 extending along the side. In the turbine stage 10, out of the combustion gas flowing into the inter-blade channel 30 along the end wall 15, the back-side recess 56 decelerates the flow velocity of the combustion gas toward the rear surface 26 of the blade 20, and the ventral surface 24 of the blade 20. The abdominal recess 54 decelerates the flow rate of the combustion gas toward. Thus, the turbine stage 10 can reduce the occurrence of a region having a high heat transfer coefficient in the end wall 15 facing the inter-blade channel 30 as compared with the turbine stage 100 that does not have the recess 50. The thermal load that the blade 20 and the platform 14 receive from the combustion gas flowing through the flow path 30 can be reduced.

  The configuration of the turbine stage for the axial-flow turbomachine according to this embodiment will be described with reference to FIG. FIG. 9 is a development view showing the shape of the end wall in the turbine stage according to the present embodiment. In the turbine stage according to the present embodiment, the recess is different from the first embodiment in that the recess has a front edge side recess extending along the back surface from the vicinity of the front edge of the blade, separately from the above-described back side recess. Details will be described below. In addition, about the structure substantially common with Example 1, the same code | symbol is attached | subjected and description is abbreviate | omitted.

  FIG. 9 is a diagram in which a platform extending in the circumferential direction (indicated by an arrow C in the drawing) of the rotating shaft of the rotor of the axial-flow turbomachine is developed linearly. Moreover, in FIG. 9, about the wing | blade, only the airfoil (aerofoil) which is the cross section is shown.

  As shown in FIG. 9, in the turbine stage 10B according to the present embodiment, the end wall, which is the wall surface of the platform (not shown) on the inter-blade channel 30 side, has a rotation axis (the axial direction is indicated by an arrow A in the drawing). Concave portions 50B that are recessed in the radial direction (shown) are arranged in the circumferential direction of the rotating shaft (indicated by arrow C in the figure).

  The concave portion 50B is substantially the same among the abdominal side concave portion 54B extending along the front edge 22 side of the abdominal surface 24 of one wing 20 and the back surface 26 of the other wing 20 provided on the abdominal surface 24 side of the wing 20. A back-side recess 56B provided in the center, and a connection recess 58 that connects the ventral-side recess 54B and the back-side recess 56B on the downstream side of the inter-blade channel 30 are provided.

  The front end 54 </ b> Ba in the axial direction A of the ventral recess 54 </ b> B is set slightly behind the front edge 22. Similarly, the front end 56 </ b> Ba in the axial direction A of the back-side recessed portion 56 </ b> B is set slightly behind the front edge 22. That is, the abdominal recess 54B and the back recess 56B according to the present embodiment are configured such that the upstream side of the inter-blade channel 30 is shorter than the abdominal recess 54 and the back recess 56 according to the first embodiment. .

  The recess 50B has a front edge side recess 57 extending along the back surface 26 from the vicinity of the front edge 22 separately from the ventral recess 54B and the back recess 56B. In the leading edge side recess 57, the downstream end 57c of the inter-blade passage 30 extends close to the front end 56Ba of the back recess 56B.

  The front edge side recess 57 configured as described above can decelerate the flow of combustion gas toward the front edge 22 of the back surface 26 as indicated by an arrow Gi in the drawing. As a result, the turbine stage 10B suppresses the formation of a region having a high heat transfer coefficient in the inter-blade channel 30, particularly the end wall 15 of the throttle 33 due to the recess 50B, and the flow of the high-speed combustion gas flows to the back surface 26. It is possible to suppress the occurrence of a turbulent flow (secondary flow) by colliding with a portion on the front edge 22 side.

  As described above, in the turbine stage 10B according to the present embodiment, the recess 50B has the front edge side recess 57 extending from the vicinity of the front edge 22 along the back surface 26, separately from the back side recess 54B. Therefore, the flow of the combustion gas toward the front edge 22 side of the back surface 26 can be decelerated, and the high-speed combustion gas flow collides with the front edge 22 side of the back surface 26 and the combustion gas flow is disturbed. This can be suppressed.

  In addition, the end wall in which the recessed part 50; 50B demonstrated in each above-mentioned Example is formed is not limited to the end wall of the platform with which the cascade of a moving blade is couple | bonded. The end wall may be a wall surface on the inter-blade flow path side of the platform to which the blades are coupled, and the recess 50; 50B is, for example, a casing (chamber) that is a platform to which the cascade of the stationary blades is coupled to the inner periphery. Or an end wall of a blade root ring in which a blade row of a stationary blade is coupled to the outer periphery.

  As described above, the present invention is useful for an axial-flow turbomachine, and is particularly suitable for a gas turbine.

FIG. 3 is a development view showing a shape of an end wall in the turbine stage according to the first embodiment. 1 is a perspective view illustrating a configuration of a turbine stage according to Embodiment 1. FIG. It is sectional drawing by the AA line of FIG. It is sectional drawing by the BB line of FIG. It is sectional drawing by CC line of FIG. It is a figure which shows the static pressure distribution in the flow path between blades of the turbine stage which concerns on a prior art. FIG. 3 is a diagram illustrating a static pressure distribution in a flow path between blades of a turbine stage according to the first embodiment. It is a figure which shows distribution of the heat transfer rate in the end wall of the turbine stage which concerns on a prior art. It is a figure which shows distribution of the heat transfer rate in the end wall of the turbine stage which concerns on Example 1. FIG. 1 is a cross-sectional view illustrating an overall configuration of a gas turbine according to Embodiment 1. FIG. FIG. 6 is a development view showing the shape of an end wall in a turbine stage according to a second embodiment.

Explanation of symbols

DESCRIPTION OF SYMBOLS 1 Gas turbine 3 Compressor 4 Combustor 5 Turbine 10, 10B Turbine stage 14 Platform 15 End wall 20 Blade 22 Front edge 24 Abdominal surface 26 Back surface 28 Rear edge 30 Inter-blade flow path 33 Restriction part 50, 50B Concave part 54, 54B Abdominal side Recess 56, 56B Back side recess 57 Front edge side recess 58 Connection recess

Claims (4)

Used in an axial-flow turbomachine, blades extending in the radial direction of the rotating shaft from a platform extending in the circumferential direction of the rotating shaft are arranged in the circumferential direction of the rotating shaft, and the inter-blade flow between adjacent blades A turbine stage in which a path is formed,
The end wall, which is the wall surface on the side of the flow path between the blades of the platform, is provided with a recess that is recessed in the radial direction of the rotating shaft.
The recess
A ventral recess extending along the front edge of the ventral surface of one of the adjacent wings;
A dorsal recess extending along the front edge side of the back surface of the other wing facing the abdominal surface;
A connection recess for connecting the back side recess and the ventral side recess to the downstream side of the flow path between the blades, and
A turbine stage characterized by comprising: The turbine stage according to claim 1 ,
The recess
A turbine stage having a front edge side recess extending along the back surface from the vicinity of the front edge separately from the back side recess. In the turbine stage according to claim 1 or 2 ,
The back-side recess and the abdomen-side recess are recessed at an equal distance in the radial direction of the rotating shaft with respect to a reference surface extending in the circumferential direction of the rotating shaft in the end wall. . A gas turbine comprising: a compressor that takes in and compresses combustion air; a combustor that combusts compressed air; and a turbine that extracts mechanical power from combustion gas,
Of the turbine stages constituting the turbine, the end wall, which is the wall surface on the side of the flow path between the blades of the platform, is provided with a recess that is recessed in the radial direction of the rotating shaft,
The recess
A ventral recess extending along the front edge of the ventral surface of one of the adjacent wings;
A dorsal recess extending along the front edge side of the back surface of the other wing facing the abdominal surface;
A connection recess for connecting the back side recess and the ventral side recess to the downstream side of the flow path between the blades, and
A gas turbine comprising:

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