JP2014101889A - Gas turbine stationary vane - Google Patents

Abstract

PROBLEM TO BE SOLVED: To provide a turbine stationary vane having effects of suppressing increase of horseshoe shaped-vortex generated near a front edge of the vane, and further suppressing secondary flow in a region held by a negative pressure surface and a pressure surface.SOLUTION: An outer peripheral side end wall inner face 10 as a face at an inner peripheral side of an outer peripheral side end wall has inward projecting shape and outward projecting shape at a negative pressure surface side of a vane profile portion 12, an apex of the inward projecting shape is positioned near a front edge 12a of the vane profile portion, an apex of the outward projecting shape is positioned near the middle between a front edge 12a and a rear edge 12b of the vane profile portion 12, the inward projecting shape is formed from an upstream side in the gas flowing direction with respect to the front edge 12a, and in a region at an upstream side in the gas flowing direction with respect to the outward projecting shape, in which the inward projecting shape is formed, a region (curved line L_end) where radius positions of the outer peripheral side end wall inner face 10 are same on a plane vertical to a turbine rotating shaft, is formed.

Description

  The present invention relates to a stationary blade of a gas turbine.

  In a blade with a large load on the blade, the flow in a cross section perpendicular to the flow of the main gas in the vicinity of the end wall, that is, the secondary flow is large near the end wall regardless of the inner peripheral side and the turbine casing side. As the secondary flow increases, the flow rate in the vicinity of the end wall decreases, and accordingly, the flow rate 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 order to prevent an increase in total pressure loss in a blade with a large load on the blade, a method of making the shape of the end wall surface a non-axisymmetric shape has been proposed. Thereby, the total pressure loss in the cascade is reduced. As an example of this, Patent Document 1 discloses a blade in which a curved surface having a pair of convex surfaces on the pressure surface side and a concave surface on the suction surface side is formed with respect to the end wall surface.

International Publication No. US 2735612A

  In order to suppress the secondary flow in the region sandwiched between the suction surface and the pressure surface, when the pressure gradient is defined as a guideline for the end wall shape, the end wall shape on the pressure surface side becomes a convex end wall shape and is negative. The end wall shape on the pressure side is determined to be concave. However, this method is expected to suppress the secondary flow in the area between the pressure surface and the suction surface, but it does not serve as a guideline for defining the endwall shape near the leading edge of the wing. An increase in the vortex cannot be suppressed, and an airfoil that is greatly affected by a horseshoe vortex has no effect.

  An object of the present invention is to provide a turbine vane having an effect of suppressing secondary flow in a region sandwiched between a suction surface and a pressure surface while suppressing an increase in a horseshoe vortex generated near the leading edge of the blade. And

An airfoil portion having a pressure surface that is concave in the chord direction and a suction surface that is convex in the chord direction;
In a gas turbine stationary blade having an outer peripheral side end wall positioned on the outer peripheral side of the airfoil portion and an inner peripheral side end wall positioned on the inner peripheral side of the airfoil portion, an inner peripheral side of the outer peripheral side end wall is provided. The inner surface of the outer end wall is an inward convex shape and an outward convex shape on the suction surface side of the airfoil portion, and the apex of the inward convex shape is located near the leading edge of the airfoil portion The apex of the outward convex shape is located in the vicinity of the middle between the front edge and the rear edge of the airfoil portion, and the inward convex shape is formed from the upstream side in the gas flow direction than the front edge. The radial position of the inner surface of the outer peripheral end wall is the same in the plane perpendicular to the turbine rotation axis in the region where the inward convex shape is formed in the previous period and in the region upstream of the outward convex shape in the gas flow direction. That the region to be And butterflies.

  ADVANTAGE OF THE INVENTION According to this invention, while suppressing the increase in the horseshoe-shaped vortex which generate | occur | produces in the front edge of a blade, the turbine stationary blade which has the effect of suppressing the secondary flow in the area | region pinched by the suction surface and the pressure surface can be provided.

It is the enlarged view which showed the stationary blade of the gas turbine. It is the figure which showed the cross-sectional shape of the airfoil part. It is explanatory drawing which showed Mach number distribution of the turbine blade surface. It is explanatory drawing which showed Mach number distribution of the turbine blade surface. It is the figure which showed the stationary blade row | line | column of the gas turbine. It is the figure which showed sectional drawing of a gas turbine. It is explanatory drawing which showed the turbine blade which is Example 1. FIG. It is explanatory drawing which showed the turbine blade which is Example 2. FIG. FIG. 6 is an explanatory view showing a turbine blade that is Embodiment 3. It is the figure which looked at the outer peripheral side end wall part inner surface from the inner peripheral side. It is the figure which looked at the inner peripheral side end wall part outer surface from the outer peripheral side. It is sectional drawing when the curved surface which forms the outer peripheral side end wall part inner surface 10 vicinity of the front edge 12a is cut | disconnected by the plane perpendicular | vertical to a turbine rotating shaft. It is sectional drawing when the curved surface which forms the inner peripheral side end wall part of the front edge 12a vicinity is cut | disconnected by the plane perpendicular | vertical to a turbine rotating shaft. It is explanatory drawing which showed the total pressure loss distribution of the turbine blade.

  Hereinafter, the present invention will be described in detail based on the illustrated embodiments.

  FIG. 6 shows a gas turbine cross-sectional view. The rotor 1 mainly includes a rotating shaft 3, a moving blade 4 disposed on the rotating shaft 3, and a moving blade (not shown) of the compressor 5. The stator 2 is mainly supported by a casing 7 and the casing 7. And a stationary blade 8 serving as a nozzle of the combustor.

  The general operation of the gas turbine configured as described above will be described. First, compressed air and fuel from the compressor 5 are supplied to the combustor 6, and these fuels burn in the combustor to generate high-temperature gas. The generated high temperature gas is blown to the moving blade 4 through the stationary blade 8 and drives the rotor 1 through the moving blade 4.

  In this case, the moving blades 4 and the stationary blades 8 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.

  FIG. 1 shows an enlarged view of the turbine stationary blade 8. A turbine stationary blade 8 shown in FIG. 1 is attached to a turbine casing 7, and has an outer peripheral side end wall portion located on the outer peripheral side, that is, on the turbine casing side with respect to the rotating shaft of the moving blade 4, and an outer peripheral side end wall portion inner surface 10. And an airfoil portion 12 extending in a direction in which the radial position becomes smaller. Further, an inner peripheral side end wall outer surface 16 is formed which forms a gas flow path surface in contact with the closed surface where the radius of the airfoil portion is smallest. The airfoil portion 12 may have a hollow portion, and the airfoil portion may be configured to cool the blade from the inside by flowing a cooling medium in the hollow portion. In FIG. 1, the inlet 9 is an inlet for the cooling medium, and the cooling medium flows in the direction of the arrow to cool the airfoil section.

  The stationary blade 8 is installed in a casing 7 that is an outer peripheral wall. As the cooling air supply source, the compressor 5 is often used, and the cooling air is introduced into the stationary blade 8 using a cooling air introduction hole provided in the casing 7. The cooled cooling air is discharged from the discharge hole 15 provided in the inner peripheral wall, and is eventually discharged to the gas path.

  FIG. 2 shows a cross-sectional shape of the airfoil portion. The airfoil portion includes a pressure surface 10b having a concave shape in the chord direction, a negative pressure surface 10a having a convex shape in the chord direction, a blade leading edge 12a, and a blade trailing edge 12b. An airfoil portion is formed so that the blade thickness gradually increases from the edge side toward the center side and gradually decreases from the middle toward the rear edge side. In some cases, the airfoil portion has hollow portions 9a and 9b, and the airfoil portion is formed so as to cool the blade from the inside by flowing a cooling medium through the hollow portion. The middle arrow in FIG. 1 indicates the flow of the cooling air, and the horizontal framed arrow indicates the flow of the hot gas, that is, the mainstream working gas.

In these FIG. 2, 12a is the leading edge, the suction surface 10a is the blade back side, the pressure surface 10b is the blade belly side, and 12b is the trailing edge. The hollow portions 9a and 9b are the air cooling chamber described above. In this case, fins are provided in the air cooling chambers 9f ₁ and 9f ₂ in front of the blades in order to improve heat conversion. As with the stationary vane of FIG. 1 after cooling, the cooling air is discharged from the discharge hole provided in the inner peripheral wall, and is eventually discharged to the gas path. This cooling structure may be convection cooling or other cooling means. What is important is the shape of the end wall of the turbine in which such cooling air is mixed.

  FIG. 3 is a diagram showing the blade surface Mach number in the blade cross section in the vicinity of the end wall on the inner peripheral side of such a turbine stationary blade. 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 inner peripheral side end wall is indicated by Ms, and the blade leading edge 12a to the blade trailing edge 12b of the pressure surface 10b in the inner peripheral end wall are indicated. The wing surface Mach number up to is indicated by Mp. As shown in FIG. 3, 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 and the blade trailing edge, and greatly decreases from the intermediate portion to the blade trailing edge. This is because gas expansion of the mainstream fluid is performed when the mainstream fluid flows from the inlet to the outlet of the cascade composed of the plurality of turbine stationary blades. M_min indicates the minimum blade surface Mach number on the pressure surface 10b. As the difference between M_max and M_min increases, the difference between the maximum pressure and the minimum pressure acting on the airfoil portion increases, and the load on the blade increases.

  In such a blade having a large load on the blade, the flow in a cross section perpendicular to the flow of the main gas in the vicinity of the end wall, that is, the secondary flow becomes large near the end wall regardless of the inner peripheral side and the turbine casing side. As the secondary flow increases, the flow rate in the vicinity of the end wall decreases, and accordingly, the flow rate in the vicinity of the average diameter increases and the blade load increases. As a result, the total pressure loss increases.

  In order to prevent such an increase in total pressure loss, a method has been proposed in which the shape of the end wall surface on the inner peripheral side and the turbine casing side is changed from an axisymmetric curved surface to a non-axisymmetric shape. Thereby, the total pressure loss in the cascade is reduced. This method is characterized in that a curved surface having a pair of convex surfaces on the pressure surface side and a concave surface on the negative pressure surface side with respect to the end wall surface is formed.

  FIG. 5 shows a turbine stationary blade row. When suppressing the secondary flow in the region sandwiched between the suction surface 10a 'and the pressure surface 10b between the stationary blades 8 arranged in the circumferential direction, pay attention to the pressure gradient as a guideline for changing the end wall shape. Can be defined. When the end wall shape is defined based on this guideline, the end wall shape on the pressure surface 10b side is determined to be a convex end wall shape, and the end wall shape on the negative pressure surface 10a side is determined to be a concave shape. Although this method has an effect of suppressing the secondary flow in the region sandwiched between the pressure surface 10b and the suction surface 10a, it does not serve as a guideline for defining the end wall shape in the vicinity of the front edge 12a, and therefore a horseshoe vortex generated from the front edge 12a. The effect is small for an airfoil that cannot suppress the increase of the horseshoe and is greatly influenced by the horseshoe vortex.

  Further, when cooling air or the like enters from the hub side upstream of the airfoil, the pressure difference between the inlet and the outlet of the hub 9 is further reduced, and the deceleration of the mainstream fluid is further increased. As a result, the total pressure loss in the blade cross section of the hub 9 is further increased.

  Hereinafter, an embodiment of a turbine vane having an effect of suppressing an increase in a horseshoe vortex generated in the vicinity of the leading edge 12a and also suppressing a secondary flow in a region sandwiched between the suction surface 10a ′ and the pressure surface 10b will be described. To do.

  Attention is paid to the stationary blade 8 shown in FIG. FIG. 7 is a perspective view of the suction surface of the airfoil portion 12 in the turbine stationary blade 8 according to the embodiment of the present invention. An arrow 13 indicates the direction of gas flow, with the leading edge 12a side being the upstream side and the trailing edge 12b being the downstream side. R is a coordinate axis representing the radial position. An outer peripheral side end wall is located on the outer peripheral side of the airfoil portion 12, and an inner peripheral end wall is located on the inner peripheral side of the airfoil portion 12. The outer peripheral side end wall portion inner surface 10 which is a surface on the inner peripheral side of the outer peripheral side end wall has an inward convex shape and an outward convex shape on the negative pressure surface side of the airfoil portion 12. Here, the outer peripheral side of the airfoil portion means a side farther from the rotor 1 than the airfoil portion 12 when the stationary blade 8 is incorporated in the gas turbine, and the inner peripheral side means the rotor 1 side. To do. Further, the outside means the outer peripheral side, and the inner means the inner peripheral side. Each convex shape only needs to be on the surface of the end wall, and the same kind of effect can be obtained regardless of whether it is in contact with or apart from the airfoil portion.

In the stationary blade 8 of the present embodiment, the apex of the inwardly convex shape on the suction surface side is located in the vicinity of the front edge.
Specifically, when the front edge of the airfoil portion in contact with the inner surface 10 on the outer peripheral side end wall is 0% and the rear edge is 100%, the vertex of the inwardly convex shape is in the range of −10% to 40%. It is formed so that it may be located in. This is because the fluid is rapidly decelerated in the region near the leading edge of the stationary blade 8 and causes vortex generation, and this is addressed. The vicinity of the front edge of the inner surface 10 of the outer peripheral side end wall portion is inwardly convex, thereby increasing the flow velocity and suppressing fluid deceleration. This effect is due to the fact that the flow velocity can be drastically increased and the generation of vortices can be suppressed by reducing the flow path by making the end wall portion inwardly convex. When the vertex of the inward convex shape is smaller than −10% and larger than 40%, the effect of suppressing the problem caused by the vortex generation near the leading edge is reduced.

  The stationary blade 8 of the present embodiment also has an apex of the outwardly convex shape on the suction surface side located in the vicinity of the middle between the front edge and the rear edge. Specifically, it is formed so as to be located within a range of 30% to 80%. This region is a region where the flow velocity increases rapidly and vortices are likely to occur. By making the convex shape outward, the flow velocity is reduced and the rapid increase of the flow velocity is suppressed. If the apex of the outwardly convex shape is smaller than 30%, the outwardly convex region becomes smaller, so the flow rate adjustment amount becomes smaller and the secondary flow suppression effect is small. On the other hand, if it is larger than 80%, a rapid increase in the flow velocity occurs downstream of the outward convex region, and the cascade performance is degraded due to shock wave loss.

  Here, the structure of the portion where the airfoil portion 12 and the end wall portion are in contact will be described. In this place, there is a rounded area called Earl. That is, the end wall portion and the airfoil portion 12 do not intersect perpendicularly. However, the size of the radius is a value ignored at the time of design. In this embodiment, a point of 0% to 100% is determined based on the contact point between the outer peripheral side end wall portion inner surface 10 and the airfoil portion 12, but this means a design contact point. Is not considered.

  Next, specific numerical values relating to the vicinity of the leading edge and the vicinity of the middle of the leading edge and the trailing edge will be described. When the apex of the inward convex shape exceeds 40%, the maximum convex amount of the convex region following the downstream side becomes the same as the radius provided on the airfoil portion and the end wall, and the effect of the convex region can be ignored. Become. For these reasons, the inwardly convex region is 40% or less. On the other hand, if the vertex of the outward convex shape is smaller than 30%, the maximum convex amount of the upstream inward convex shape becomes larger than 80%, and the maximum convex amount of the outward convex shape becomes about the same as the radius. . Therefore, the outward convex region is set to be 30% or more and 80% or less.

  As described above, the stationary blade 8 of the present embodiment has a radial position that decreases from the upstream side of the gas flow in the vicinity of the negative pressure portion of the inner surface 10 of the outer peripheral side end wall portion that is the end wall on the turbine casing 7 side. An outwardly convex shape is formed, and the radial position increases from there toward the downstream, thereby forming an outwardly convex shape. By forming the stationary blade 8 in such a shape, it is possible to smoothly change the speed change while suppressing rapid deceleration and acceleration of the flow in the main flow direction indicated by the arrow 13, and a suitable stationary blade 8 is provided. it can. Each convex shape only needs to be on the end wall, and the same kind of effect can be obtained regardless of whether it is in contact with or apart from the airfoil portion 12.

  In the gas turbine configured as described above, the main flow fluid that flows in toward the turbine stationary blade 8 flows in from the blade leading edge 12a, flows along the airfoil portion, and flows out from the blade trailing edge 12b. By adopting this end wall shape, the secondary flow is suppressed, the deceleration of the flow near the end wall on the outer peripheral side of the mainstream fluid flowing along the airfoil negative pressure surface 10a is suppressed, and the airfoil part negative of the stationary blade 8 is suppressed. The decrease in Mach number on the pressure surface 10a is also reduced. As a result, it is possible to reduce the total pressure loss in the blade cross section of the airfoil portion suction surface 10a of the stationary blade 8. Even when the aerodynamic load is high or when the mixed cooling flow rate is changed, an increase in the total pressure loss in the blade cross section of the hub can be suppressed.

  In addition, the outer peripheral side end wall part inner surface 10 is a surface which forms a gas flow path surface. On the outer peripheral side of the end wall, there is an outer peripheral side end wall portion outer surface 10 ′ that is paired with the outer peripheral side end wall portion inner surface 10. The outer peripheral end wall thickness, which is the distance between the outer peripheral end wall portion outer surface 10 ′ and the outer peripheral end wall portion inner surface 10, may or may not be constant.

  FIG. 8 is a perspective view of an airfoil suction surface 10a of a turbine vane 8 according to another embodiment of the present invention. The same parts as in FIG. 7 are omitted, and only the differences will be described. The inner peripheral side end wall portion outer surface 16 that is the outer peripheral side surface of the inner peripheral side end wall has an outwardly convex shape and an inwardly convex shape on the negative pressure surface side of the airfoil portion 12.

In the stationary blade 8 of the present embodiment, the apex of the outwardly convex shape on the suction surface side is located in the vicinity of the front edge.
Specifically, when the leading edge of the airfoil portion in contact with the outer surface 16 on the inner peripheral side end wall is 0% and the trailing edge is 100%, the apex of the outwardly convex shape is in the range of −10% to 40%. It is formed so as to be located inside. This is because the fluid is rapidly decelerated in the region near the leading edge of the stationary blade 8 and causes vortex generation, and this is addressed. The vicinity of the front edge of the outer surface 16 on the inner peripheral side end wall portion is formed in an outwardly convex shape, thereby increasing the flow velocity and suppressing the deceleration of the fluid. This effect is due to the fact that the flow velocity can be rapidly increased by reducing the flow path by making the end wall portion outwardly convex, and the generation of vortices can be suppressed. When the apex of the outward shape is smaller than −10% and larger than 40%, the effect of suppressing the problem caused by the vortex generation near the leading edge is reduced.

  The stationary blade 8 of the present embodiment also has an inwardly convex vertex on the suction surface side located in the vicinity of the middle between the front edge and the rear edge. Specifically, it is formed so as to be located within a range of 30% to 80%. This region is a region where the flow velocity increases rapidly and vortices are likely to occur. By making it inwardly convex, the flow velocity is reduced and the rapid increase of the flow velocity is suppressed. When the vertex of the inwardly convex shape is smaller than 30%, the inwardly convex region becomes smaller, so the flow rate adjustment amount becomes smaller and the secondary flow suppression effect is small. On the other hand, if it is larger than 80%, a rapid increase in flow velocity occurs downstream of the inwardly convex region, and the cascade performance deteriorates due to shock wave loss. The apex of the outward convex shape on the suction surface side and the apex of the inward convex shape are based on the aerodynamic design conditions of the turbine to be designed to suppress rapid deceleration and acceleration of the flow in the main flow direction indicated by arrow 13. It is selected so as to be optimal among the above conditions so that the speed change smoothly changes.

  Next, specific numerical values relating to the vicinity of the leading edge and the vicinity of the middle of the leading edge and the trailing edge will be described. When the apex of the outward convex shape exceeds 40%, the maximum convex amount of the inward convex region following the downstream side becomes the same as the radius provided on the airfoil and end wall, and the effect of the convex region is ignored. It will be to the extent possible. For these reasons, the outwardly convex region is 40% or less. On the other hand, when the vertex of the inward convex shape is smaller than 30%, the maximum convex amount of the upstream outward convex region becomes larger than 80%, and the maximum convex amount of the inward convex shape becomes approximately the same as the radius. Therefore, the inwardly convex region is set to be 30% or more and 80% or less.

  As described above, the stationary blade 8 of this embodiment has a radial position that increases from the upstream side of the gas flow in the vicinity of the negative pressure portion of the inner peripheral side end wall portion outer surface 16 that is the end wall on the rotor 1 side. An inwardly convex shape is formed, and from there, the radial position becomes smaller and the inwardly convex shape is formed.

  In the gas turbine configured as described above, the main flow fluid that flows in toward the turbine stationary blade 8 flows in from the blade leading edge 12a, flows along the airfoil portion 12, and flows out from the blade trailing edge 12b. The outward convex region and the inward convex region are set in the above direction in the flow direction, so that the speed change becomes moderate and the secondary flow loss is suppressed. As a result, the total pressure loss in the blade cross section of the hub of the airfoil portion 12 can be reduced.

  In addition, the inner peripheral side end wall part outer surface 16 is a surface which forms a gas flow path surface. On the inner peripheral side of the end wall, there is an inner peripheral end wall inner surface 16 ′ that is paired with the inner peripheral end wall outer surface 16. The thickness of the inner peripheral end wall, which is the distance between the inner peripheral end wall inner surface 16 ′ and the inner peripheral end wall outer surface 16, may or may not be constant.

  FIG. 9 is a perspective view of a suction surface of an airfoil portion 12 of a turbine vane according to another embodiment of the present invention. The description of the parts common to FIGS. 7 and 8 is omitted. This embodiment is a combination of Example 1 and Example 2. That is, the outward convex shape of the inner peripheral side end wall portion outer surface 16 of the stationary blade 8 of the first embodiment is located in the vicinity of the front edge 12a, and the inward convex shape vertex of the inner peripheral side end wall portion outer surface 16 is the blade. It is located near the middle between the front edge and the rear edge of the mold part 12. The vane 8 of the present embodiment can provide the more suitable vane by being able to enjoy the advantages of both examples.

  Next, FIGS. 10-13 which looked at the stationary blade of each Example from another angle are demonstrated.

  FIG. 10 shows a view of the outer peripheral side end wall portion inner surface 10 as viewed from the inner peripheral side. The vertical dotted line portion is formed so that the radial position is low, and the horizontal dotted line portion is formed so that the radial position is high. 13a is a negative pressure side flow of the turbine casing side end wall portion, and 13b is a pressure side flow of the turbine casing side end wall portion.

  On the suction surface side of the inner surface 10 of the outer peripheral side end wall portion, in the flow direction 13a, a transition is made from a convex region where the radial position of the rotor 1 is reduced near the front edge to a convex region where the radial position is increased. ing. Further, in the flow direction 13b on the pressure surface side, a transition is made from a convex region in which the radial position becomes smaller near the front edge to a convex shape in which the radial position becomes larger. The point to be noted is that the concave surface and the convex surface are not paired on the pressure surface side and the suction surface side of the end wall part, but the concave surface and the convex surface are paired in the flow direction on both the suction surface side and the pressure surface side. It is.

  FIG. 12 is a cross-sectional view of the curved surface forming the outer peripheral end wall portion inner surface 10 in the vicinity of the front edge 12a in FIG. 10 taken along a plane perpendicular to the turbine rotation axis. The cross section of this curved surface is a curve L_end, the intersection with the suction surface of the airfoil portion 12 is a point C, and the intersection with the pressure surface is a point D. The curve L_end extends smoothly from the intersection C to the intersection D. The radial position of the curve L_end is the same. The radial position of the intersection point C, the intersection point D, and the shape of the curved surface L_end are selected so as to be optimal based on the aerodynamic design conditions of the turbine to be designed.

  The curve L_end has the same radial position in the vicinity of the turbine casing side end wall in the vicinity of the leading edge 12a in FIG. 10, but the condition that the radial position is the same is not set in the entire region. If the condition that the radial positions are the same is set in the entire region, a shock wave that greatly affects the increase in the total pressure loss of the turbine blades is generated. As a result, the pressure ratio condition at the inlet / outlet of the blade is reduced, and the performance of the turbine blade is degraded.

  FIG. 11 shows a view of the outer peripheral wall 16 on the inner peripheral side as viewed from the outer peripheral side. The vertical dotted line portion is formed so that the radial position is high, and the horizontal oblique line portion is formed so that the radial position is small. 13a is a flow on the negative pressure side of the end wall portion, and 13b is a flow on the pressure surface side of the end wall portion. At this time, on the suction surface side of the outer surface on the inner peripheral side end wall portion, a convex region in which the radial position is decreased from a convex region in which the radial position of the rotor 1 is increased near the front edge in the flow direction 13a. Has moved to. Further, in the flow direction 13b on the pressure surface side, a transition is made from a convex region in which the radial position is increased near the leading edge to a convex shape in which the radial position is decreased. The point to be noted is that the concave surface and the convex surface are not paired on the pressure surface side and the suction surface side of the end wall part, but the concave surface and the convex surface are paired in the flow direction on both the suction surface side and the pressure surface side. It is.

  FIG. 13 is a cross-sectional view of the curved surface forming the inner peripheral side end wall portion in the vicinity of the front edge 12a in FIG. 11 taken along a plane perpendicular to the turbine rotation axis. A cross section of this curved surface is a curve L_end, an intersection with the suction surface of the airfoil portion is a point C, and an intersection with the pressure surface is a point D. The curve L_end extends smoothly from the intersection C to the intersection D. The radial position of the curve L_end is the same. The radial positions of the intersection point C and the intersection point D and the upper surface shape contour of the curved surface 10 are selected so as to be optimal based on the aerodynamic design conditions of the turbine to be designed.

  The curve L_end has the same radial position in the vicinity of the outer peripheral end wall 16 near the front edge 12a in FIG. 10, but the condition that the radial position is the same is not set in the entire region. If the condition that the radial positions are the same is set in the entire region, a shock wave that greatly affects the increase in the total pressure loss of the turbine blades is generated. As a result, the pressure ratio condition at the inlet / outlet of the blade is reduced, and the performance of the turbine blade is degraded.

  FIG. 14 is a diagram showing the total pressure loss of the blade cross section over the airfoil portion in the vertical direction. FIG. 14 shows a comparison between the comparative example in which the end wall has no local unevenness and the above-described embodiment. In the comparative example, as shown by the solid line, particularly significant blade section total pressure loss is observed in the end wall. However, in this embodiment, as shown by the solid line, the end wall on the inner peripheral side and the end on the turbine casing side are shown. The total pressure loss in the blade cross section of the wall is reduced, and a more uniform total pressure loss is achieved in the vertical direction of the airfoil portion. This means that more uniform expansion work is achieved in the up-down direction of the airfoil, thereby improving turbine efficiency and reducing gas turbine fuel consumption.

DESCRIPTION OF SYMBOLS 1 Rotor 2 Stator 3 Rotating shaft 4 Rotor blade 5 Compressor 6 Combustor 7 Casing 8 Stator blade 9h Fin 9a, 9b Hollow part 9c Pressure surface 9f Air cooling chamber 10 Outer end wall inner surface 10a Negative pressure surface 10b Pressure surface 12 Blade Mold part 12a Front edge 12b Rear edge 13 Arrow 15 Discharge hole 16 Inner peripheral side end wall outer surface

Claims (5)

An airfoil portion having a pressure surface that is concave in the chord direction and a suction surface that is convex in the chord direction;
An outer peripheral side wall located on the outer peripheral side of the airfoil, and
In the gas turbine stationary blade having the inner peripheral side end wall located on the inner peripheral side of the airfoil part,
The outer peripheral side end wall inner surface which is the inner peripheral side surface of the outer peripheral side end wall has an inward convex shape and an outward convex shape on the suction surface side of the airfoil portion,
The vertex of the inwardly convex shape is located near the leading edge of the airfoil,
The apex of the outward convex shape is located in the vicinity of the middle of the front edge and the rear edge of the airfoil part,
The inward convex shape is formed from the upstream side in the gas flow direction from the front edge,
In the region where the inward convex shape is formed and upstream of the outward convex shape in the gas flow direction, the radial position of the inner surface of the outer peripheral end wall is the same in the plane perpendicular to the turbine rotation axis. A gas turbine vane characterized in that a region is formed. An airfoil portion having a pressure surface that is concave in the chord direction and a suction surface that is convex in the chord direction;
An outer peripheral side wall located on the outer peripheral side of the airfoil, and
In the gas turbine stationary blade having the inner peripheral side end wall located on the inner peripheral side of the airfoil part,
The inner peripheral side end wall outer surface which is the outer peripheral side surface of the inner peripheral side end wall has an outwardly convex shape and an inwardly convex shape on the negative pressure surface side of the airfoil portion,
The apex of the outwardly convex shape is located in the vicinity of the leading edge of the airfoil portion, and the apex of the inwardly convex shape is located in the vicinity of the middle between the front edge and the trailing edge of the airfoil portion,
The outward convex shape is formed from the upstream side in the gas flow direction from the front edge,
In the region where the outward convex shape is formed and upstream of the inward convex shape in the gas flow direction, the radial position of the outer surface of the inner end wall is the same in the plane perpendicular to the turbine rotation axis. A gas turbine stationary blade characterized in that a region to be formed is formed. The gas turbine stationary blade according to claim 1,
The apex of the outwardly convex shape on the outer surface of the inner peripheral side end wall is located in the vicinity of the leading edge of the airfoil portion, and the apex of the inwardly convex shape of the outer surface of the inner peripheral side endwall is the front edge of the airfoil portion. A gas turbine stationary blade, characterized by being positioned in the vicinity of the middle of the trailing edge. In any one of Claim 1 to 3,
The vicinity of the leading edge means that the contact between the end wall and the leading edge of the airfoil is 0%, and the contact between the end wall and the trailing edge of the airfoil is 100%. % Of a gas turbine stationary blade, characterized by being within a range of% or less. In any one of Claim 1-4,
The intermediate vicinity means 30% or more and 80% or less when the contact between the end wall and the leading edge of the airfoil is 0% and the contact between the end wall and the trailing edge of the airfoil is 100%. A gas turbine stationary blade characterized by being in the range of.

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