Gas turbine

Abstract

A turbine is provided with: a turbine rotor; a turbine casing; a plurality of turbine rotor blade rows; a plurality of turbine stator blade rows; and a diffuser provided on the other side of the axial direction of the last stage rotor blade row and forming an exhaust flow path in which exhaust gas flows from the one side of the axial direction toward the other side, the diffuser having: an inner cylinder; an outer cylinder forming the exhaust flow path with the inner cylinder; and a strut connecting the inner cylinder and the outer cylinder in the radial direction, the outer cylinder having a first inclined surface extending from the inner side to the outer side in the radial direction as it goes from the inlet of the exhaust flow path toward the other side of the axial direction, the first inclined surface being at an angle of 16° or more and 24° or less with respect to the axis, and in the turbine rotor blades of the last stage rotor blade row, the nozzle width of the end portion on the outer side in the radial direction with respect to the axis is set to be greater than the nozzle width of the middle portion in the radial direction.

Description

Technical Field

[0001] This invention relates to a turbine and a gas turbine.

[0002] This application claims priority based on Japanese Patent Application No. 2021-050511, filed on March 24, 2021, the contents of which are incorporated herein by reference. Background Technology

[0003] A gas turbine mainly comprises: a compressor that generates compressed air; a combustor that generates high-temperature, high-pressure combustion gases by mixing and burning fuel with compressed air; a turbine driven by the combustion gases; and a cylindrical diffuser that guides the exhaust gas from the turbine. As illustrated in Patent Document 1 below, the diffuser comprises: an inner cylinder extending along the axis of the gas turbine; an outer cylinder forming an exhaust flow path between the inner cylinder and the outer cylinder by being disposed on the outer periphery of the inner cylinder; and a strut connecting the inner cylinder and the outer cylinder. In the diffuser described in Patent Document 1, the outer cylinder gradually increases in diameter towards the downstream side. That is, in a cross-section including the axis, the inner circumferential surface of the outer cylinder is inclined relative to the axis. As a result, the flow of exhaust gas from the turbine slows down and restores static pressure as it passes through the diffuser.

[0004] It is known that pressure losses occur in the exhaust flow as it passes through the diffuser. Most of these pressure losses are due to the aforementioned support being exposed to the exhaust flow. To suppress these pressure losses in the support and improve diffuser performance, it is crucial to reduce the exhaust velocity upstream of the support. Therefore, it is necessary to maximize the inclination angle of the inner circumferential surface of the outer cylinder.

[0005] Previous technical documents

[0006] Patent documents

[0007] Patent Document 1: Japanese Patent No. 6018368 Summary of the Invention

[0008] The technical problem to be solved by the invention

[0009] However, if the inclination angle of the inner circumferential surface of the outer cylinder is too large, there is a problem of flow separation from that inner circumferential surface, generating vortices. The generation of vortices will impair the aerodynamic performance of the diffuser. Consequently, this may affect the performance of the turbine.

[0010] The present invention was made to solve the above-mentioned problems, and its purpose is to provide a turbine and gas turbine that further improves performance by reducing pressure loss.

[0011] means for solving technical problems

[0012] To address the aforementioned issues, the turbine of the present invention comprises: a turbine rotor extending along an axis and rotatable about the axis; a turbine housing covering the turbine rotor from its outer peripheral side; a plurality of turbine moving blade rows having a plurality of turbine moving blades arranged circumferentially along the axis on the outer peripheral surface of the turbine rotor, and arranged along the axial direction; a plurality of turbine stationary blade rows having a plurality of turbine stationary blades arranged circumferentially on the inner peripheral surface of the turbine housing in an adjacent manner to the turbine moving blades on one side in the axial direction, and arranged along the axial direction; and a diffuser disposed on the other side of the axial direction of the last-stage moving blade row on the other side of the axial direction in the plurality of turbine moving blade rows, and forming exhaust gas from one side of the axial direction to the other side. The diffuser has a flowing exhaust path, comprising: an inner cylinder extending along the axis; an outer cylinder covering the inner cylinder from the outer periphery and forming the exhaust path between the outer cylinder and the inner cylinder; and a plurality of struts disposed at the midpoint of the exhaust path, connecting the inner cylinder and the outer cylinder radially and arranged circumferentially, the outer cylinder having a first inclined surface extending radially inward from the axis as the inlet of the exhaust path on one side of the axis faces the other side, the first inclined surface forming an angle of 16° or more and 24° or less relative to the axis in a cross-section including the axis, wherein the nozzle width at the radially outer end of the turbine blades in the last stage blade row is set to be greater than the nozzle width at the radially middle portion.

[0013] Invention Effects

[0014] According to the present invention, a turbine and gas turbine can be provided that further improves performance by reducing pressure loss. Attached Figure Description

[0015] Figure 1 This is a cross-sectional view showing the schematic structure of a gas turbine according to an embodiment of the present invention.

[0016] Figure 2 This is an enlarged cross-sectional view of the main part of the gas turbine involved in the embodiments of the present invention.

[0017] Figure 3 This is a perspective view showing the structure of the final stage moving blade array according to an embodiment of the present invention.

[0018] Figure 4 This is a diagram showing the final stage moving blade array according to the embodiment of the present invention viewed radially, and is an explanatory diagram showing the nozzle width and discharge angle at the tip-side end and the hub-side end.

[0019] Figure 5This is a diagram showing the final stage moving blade array according to the embodiment of the present invention viewed radially, and an explanatory diagram showing the nozzle width and discharge angle in the middle section.

[0020] Figure 6 It is a curve representing the relative outflow angle of the moving blades in the height direction of the last stage moving blade row.

[0021] Figure 7 It is a graph representing the absolute total pressure at the outlet of the last stage moving blade row in the height direction. Detailed Implementation

[0022] (Structure of a gas turbine)

[0023] The following is for reference. Figures 1 to 7 The gas turbine 10 according to the embodiments of the present invention will be described.

[0024] like Figure 1 As shown, the gas turbine 10 includes a compressor 20, a burner 30, a turbine 40, and a diffuser 60. The compressor 20 compresses air A drawn in from the outside to generate high-pressure compressed air. The burner 30 mixes fuel F with the compressed air and combusts it to generate high-temperature, high-pressure combustion gas G. The turbine 40 is driven by the rotation of the combustion gas G.

[0025] The compressor 20 has a compressor rotor 21, a compressor housing 25, and multiple rows of compressor stationary blades 26. The compressor rotor 21 has a compressor rotating shaft 22 and multiple rows of compressor moving blades 23. The compressor rotating shaft 22 extends along an axis Ac and is rotatable about this axis Ac. Multiple rows of compressor moving blades 23 are arranged at intervals along the outer circumferential surface of the compressor rotating shaft 22 in the direction of the axis Ac. Although not shown in detail, each row of compressor moving blades 23 has multiple compressor moving blades arranged circumferentially along the outer circumferential surface of the compressor rotating shaft 22.

[0026] The compressor housing 25 is cylindrical, covering the compressor rotor 21 from the outer periphery. On the inner periphery of the compressor rotor 21, a plurality of compressor stationary blade rows 26 are arranged along the axis Ac. The aforementioned compressor moving blade rows 23 and compressor stationary blade rows 26 are arranged alternately along the axis Ac. More specifically, one compressor stationary blade row 26 is provided on one side of one compressor moving blade row 23 along the axis Ac.

[0027] An intermediate chamber 16 is connected to the other side of the compressor housing 25 along the Ac direction. The burner 30 is disposed in this intermediate chamber 16.

[0028] The turbine 40 includes a turbine rotor 41, a turbine housing 45, and multiple rows of stationary turbine blades 46. The turbine rotor 41 has a turbine rotating shaft 42 and multiple rows of moving turbine blades 43. The turbine rotating shaft 42 extends along an axis Ac and is rotatable about this axis Ac. Multiple rows of moving turbine blades 43 are arranged at intervals along the outer circumferential surface of the turbine rotating shaft 42 in the direction of the axis Ac. Although not shown in detail, each row of moving turbine blades 43 has multiple moving turbine blades arranged circumferentially along the outer circumferential surface of the turbine rotating shaft 42. Furthermore, among the multiple rows of moving turbine blades 43 arranged along the axis Ac, the row of moving turbine blades 43 on the far side in the direction of the axis Ac is designated as the last-stage moving blade row 43A. In the following description, the turbine blades included in this last-stage moving blade row 43A are referred to as the last-stage moving blade row 50.

[0029] The compressor rotor 21 and turbine rotor 41 described above are coaxially connected along axis Ac to form gas turbine rotor 11. The compressor housing 25, intermediate chamber 16, and turbine housing 45 are coaxially connected along axis Ac to form gas turbine housing 15. Gas turbine rotor 11 can rotate integrally around axis Ac inside gas turbine housing 15. Furthermore, in the following description, when viewed from turbine 40, the side where compressor 20 is located (i.e., the side in the direction of axis Ac) is sometimes simply referred to as the "upstream side," and the opposite side (i.e., the other side in the direction of axis Ac) is sometimes simply referred to as the "downstream side."

[0030] The diffuser 60 is provided to reduce the flow rate of the gas (exhaust gas) discharged from the turbine 40 and restore static pressure. The diffuser 60 is connected to the downstream side of the turbine housing 45. The diffuser 60 has an inner cylinder 62, an outer cylinder 61, a plurality of supports 63, and a plurality of manholes 64. The inner cylinder 62 extends along the axis Ac. Inside the inner cylinder 62, a bearing assembly 80 (described later) for rotatably supporting the gas turbine rotor 11 is housed. In this embodiment, as an example, the outer diameter of the inner cylinder 62 is constant from the upstream side to the downstream side. Alternatively, a structure in which the outer diameter of the inner cylinder 62 gradually decreases from the upstream side to the downstream side can also be adopted.

[0031] The outer cylinder 61 is cylindrical, covering the inner cylinder 62 from the outer periphery. The space between the outer cylinder 61 and the inner cylinder 62 is configured as an exhaust flow path E through which gas discharged from the turbine 40 flows. The inner diameter of the outer cylinder 61 gradually increases from the upstream side to the downstream side. Therefore, the cross-sectional area of ​​the exhaust flow path E gradually increases from the downstream side.

[0032] The outer cylinder 61 and the inner cylinder 62 are radially connected by a support 63. That is, the support 63 supports the outer cylinder 61 radially inward relative to the inner cylinder 62. The support 63 is positioned midway along the exhaust flow path E in the direction of axis Ac. The detailed configuration of the support 63 will be described later. Multiple supports 63 are arranged at intervals in the circumferential direction. Each support 63 extends radially between the inner circumferential surface of the outer cylinder 61 and the outer circumferential surface of the inner cylinder 62. Furthermore, although not shown in detail, it is preferable that the support 63 has a streamlined cross-sectional shape from the upstream side to the downstream side when viewed radially.

[0033] Downstream of support 63, manholes 64 are spaced apart along axis Ac. Manholes 64 extend radially between outer cylinder 61 and inner cylinder 62. Multiple manholes 64 are arranged at circumferential intervals. Various pipes or wiring are accommodated inside the manholes 64.

[0034] (Detailed structure of the last stage moving blade array and diffuser)

[0035] Next, refer to Figure 2 The detailed structure of the aforementioned final-stage moving blade row 50 and diffuser 60 is described below. Figure 2 As shown, the final stage moving blade row 50 has a disk 70 and a moving blade body 50H. The disk 70 is mounted on the turbine rotating shaft 42. The disk 70 is disc-shaped with the axis Ac as its center. The moving blade body 50H is disposed on the outer peripheral surface 70A of the disk 70. The moving blade body 50H extends radially outward from the outer peripheral surface 70A.

[0036] Detailed descriptions will follow, but when viewed radially, the moving blade body 50H has a blade-like cross-sectional shape. The upstream-facing end edge of the moving blade body 50H is designated as the leading edge 50A. The downstream-facing end edge of the moving blade body 50H is designated as the trailing edge 50B. The radially outward-facing end face of the moving blade body 50H is designated as the tip-side end face 50C. The tip-side end face 50C extends radially inward and outward from the upstream side towards the downstream side. That is, in a cross-section including the axis Ac, the tip-side end face 50C is inclined at an angle θ1 relative to the axis Ac. This angle θ1 is appropriately set within a range of 20° to 25°. The tip-side end face 50C is radially spaced from the inner circumferential surface 45A of the turbine housing 45. The inner diameter of the inner circumferential surface 45A gradually increases from the upstream side towards the downstream side.

[0037] Furthermore, the radially inner end of the moving blade body 50H is designated as a hub-side end face 50D. The hub-side end face 50D abuts against the outer peripheral surface 70A of the wheel disk 70. Although not shown in detail, a blade root with serrated protrusions and concavities is formed on the radially inner side of the hub-side end face 50D. A groove corresponding to this protrusion and concavity shape is formed in the wheel disk 70. The moving blade body 50H is supported in a way that prevents it from detaching by engaging with the inner surface of the groove in the protrusion and concavity shape formed on the blade root.

[0038] The inner cylinder 62 of the diffuser 60 covers the shaft end of the turbine rotating shaft 42 from the outer peripheral side. A bearing assembly 80 is disposed inside the inner cylinder 62. The bearing assembly 80 rotatably supports the turbine rotating shaft 42. Specifically, a journal bearing is exemplified as the bearing assembly 80 located in this position. The journal bearing supports the radial load generated by the turbine rotating shaft 42. The outer peripheral surface of the inner cylinder 62 is designated as the outer peripheral surface 62A. This outer peripheral surface 62A is located at the same radial position as the outer peripheral surface 70A of the aforementioned disc 70. Furthermore, the term "same" here means substantially the same, and allows for design tolerances or manufacturing errors.

[0039] The inner circumferential surface of the outer cylinder 61 is formed by a first inclined surface 61A and a second inclined surface 61B. The first inclined surface 61A is connected to the downstream side of the inner circumferential surface 45A of the turbine housing 45. The first inclined surface 61A extends radially inward from the upstream side to the downstream side. That is, in a cross-section including the axis Ac, the first inclined surface 61A is inclined at an angle θ2 relative to the axis Ac. This angle θ2 is appropriately determined in the range of 16° or more and 24° or less. Therefore, the difference between the angle θ1 formed by the blade tip side end face 50C of the aforementioned moving blade body 50H with respect to the axis Ac and the angle θ2 formed by the first inclined surface 61A with respect to the axis Ac is in the range of 0° or more and 5° or less.

[0040] The second inclined surface 61B is connected to the downstream side of the first inclined surface 61A. Like the first inclined surface 61A, the second inclined surface 61B extends radially inward from the upstream side to the downstream side. That is, the second inclined surface 61B is inclined relative to the axis Ac. The angle formed by the second inclined surface 61B relative to the axis Ac is smaller than the angle θ2 formed by the first inclined surface 61A relative to the axis Ac. More specifically, this angle is preferably set to approximately 8°.

[0041] The upstream end edge of the aforementioned support column 63 is designated as the leading edge 63A, and the downstream end edge is designated as the trailing edge 63B. The leading edge 63A is located on the first inclined surface 61A. On the other hand, the trailing edge 63B is located on the second inclined surface 61B. That is, the support column 63 is configured to span the first inclined surface 61A and the second inclined surface 61B in the direction of the axis Ac.

[0042] Here, the cross-sectional area (viewed from the direction of axis Ac) of the exhaust flow path E at the inlet (upstream end) of the diffuser 60 is set as S1. Furthermore, the cross-sectional area (viewed from the direction of axis Ac) of the exhaust flow path E in the leading edge 63A of the support column is set as S2. That is, cross-sectional areas S1 and S2 refer to the areas of the annular region surrounded by the outer peripheral surface 62A of the inner cylinder 62 and the inner peripheral surface (first inclined surface 61A) of the outer cylinder 61. In this case, the ratio (area ratio) of cross-sectional area S1 to cross-sectional area S2 is preferably in the range of 1.25 or more and 1.40 or less. More preferably, this area ratio is in the range of 1.28 or more and 1.37 or less. Most preferably, this area ratio is set to 1.30. Furthermore, when achieving the above-mentioned ratio of cross-sectional areas, the reduction in area caused by the rounded corner portion formed in the mounting portion of the support column 63 can be disregarded.

[0043] (More detailed structure of the last stage moving blade row)

[0044] Next, refer to Figures 3 to 6 A more detailed explanation of the structure of the last-stage moving blade row 50 (moving blade body 50H) is provided. For example... Figure 3 As shown, when viewed radially, the moving blade body 50H has a blade-like cross-sectional shape. That is, the moving blade body 50H bends circumferentially from the leading edge 50A toward the trailing edge 50B. The circumferentially facing surface of the moving blade body 50H (i.e., the rearward side facing the rotation direction of the turbine shaft 42) is designated as a positive pressure surface 50P. The positive pressure surface 50P is concave on the other circumferential side. The circumferentially facing surface of the moving blade body 50H (i.e., the forward side facing the rotation direction of the turbine shaft 42) is designated as a negative pressure surface 50N. The negative pressure surface 50N protrudes in a curved shape on the other circumferential side.

[0045] Furthermore, in the main body 50H of the moving blade, the cross-sectional shape differs at the tip-side end face 50C, the hub-side end face 50D, and the radially intermediate position (middle part 50M). Here, as... Figure 4 As shown, the nozzle width in the blade tip side end face 50C and the hub side end face 50D is set to A1. Furthermore, the outflow angle in the blade tip side end face 50C and the hub side end face 50D is set to θ3. Moreover, as... Figure 5 As shown, the nozzle width in the middle section 50M is set to A2. Furthermore, the outflow angle in the middle section 50M is set to θ4.

[0046] At this point, as shown in these figures, the nozzle width A1 in the blade tip side end face 50C and the hub side end face 50D is set to be greater than the nozzle width A2 in the middle part 50M. Furthermore, the nozzle width referred to here is the flow path width at the position (nozzle position) where the separation distance between a pair of adjacent moving blade bodies 50H in the circumferential direction is minimal. The nozzle width gradually decreases from the blade tip side end face 50C towards the middle part 50M, and gradually increases from the middle part 50M towards the hub side end face 50D.

[0047] Furthermore, the outflow angle θ3 in the blade tip side end face 50C and the hub side end face 50D is set to be smaller than the outflow angle θ4 in the middle part 50M. Additionally, the outflow angle referred to here is the angle relative to the axis Ac formed by the flow between a pair of adjacent moving blade bodies 50H. The outflow angle gradually increases from the blade tip side end face 50C towards the middle part 50M, and gradually decreases from the middle part 50M towards the hub side end face 50D.

[0048] like Figure 6 As shown by the double-dotted line, in conventional moving blades, the outflow angle is set to gradually decrease from the blade tip side towards the hub side. In contrast, in the moving blade body 50H of this embodiment, as shown by the solid line, the outflow angle is set to gradually increase from the blade tip side towards the hub side and then gradually decrease towards the hub side. Thus, in the moving blade body 50H, the outflow angle is smaller at both the blade tip side and the hub side than at the middle section 50M. That is, the nozzle width is greater at both the blade tip side and the hub side than at the middle section 50M.

[0049] (Effects)

[0050] Next, the operation of the gas turbine 10 according to this embodiment will be described. When driving the gas turbine 10, a rotational force is first applied to the compressor rotor 21 by an external drive source. As a result, the compressor rotor 21 rotates about the axis Ac, and external air A is drawn into the compressor 20. The air A drawn into the compressor 20 comes into contact with the compressor stationary blade row 26 and the compressor moving blade row 23 as it flows from the upstream side to the downstream side through the compressor housing 25. The compressor stationary blade row 26 changes the flow direction of the air A, optimizing the inflow angle of the compressor moving blade row 23 to the downstream side. Moreover, the air A gradually increases in pressure by being compressed by the compressor moving blade row 23, thus becoming compressed air. The compressed air is guided to the combustor 30 located downstream of the compressor 20. In the combustor 30, the compressed air is mixed with fuel F and burned to generate combustion gas G. The combustion gas G is guided to the turbine 40 located downstream of the combustor 30.

[0051] In turbine 40, combustion gas G comes into contact with turbine stationary blade row 46 and turbine moving blade row 43. Turbine stationary blade row 46 alters the flow direction of combustion gas G, facilitating the inflow angle of the downstream turbine moving blade row 43. Furthermore, as combustion gas G flows around turbine moving blade row 43, it imparts rotational energy to turbine rotor 41 via turbine moving blade row 43. Consequently, gas turbine rotor 11 rotates about axis Ac. Gas discharged from turbine 40, after regaining static pressure while flowing through diffuser 60, is guided to other external equipment (not shown).

[0052] Here, it is known that pressure loss occurs in the exhaust flow when passing through the diffuser 60. Most of this pressure loss is generated through the aforementioned support 63 being exposed to the exhaust flow. To suppress the pressure loss in the support 63 and improve the performance of the diffuser 60, it is crucial to reduce the exhaust velocity upstream of the support 63. Therefore, it is necessary to maximize the inclination angle of the inner circumferential surface of the outer cylinder 61.

[0053] However, if the inclination angle of the inner circumferential surface of the outer cylinder 61 is too large, the following problem arises: the flow cannot fully follow the inner circumferential surface, resulting in flow stripping and the generation of vortices. The generation of vortices will impair the aerodynamic performance (static pressure recovery) of the diffuser 60. Consequently, this may affect the performance of the gas turbine 10.

[0054] Therefore, the structure described above is adopted in this embodiment. According to the above structure, the angle between the first inclined surface 61A of the outer cylinder 61 and the axis Ac is set to be 16° or more and 24° or less, which is larger than that of a conventional diffuser. This allows for a further reduction in the flow velocity of the exhaust gas flowing through the exhaust path E in a region further upstream of the support column 63. As a result, the flow of exhaust gas around the support column 63 is less affected by the influence generated by the support column 63. That is, pressure loss generated by the support column 63 can be further reduced.

[0055] Furthermore, in the last-stage moving blade row 50, the nozzle width A1 at the radially outer end (tip side) is greater than the nozzle width A2 at the middle part 50M. As a result, the total pressure at the radially outer end is higher. Therefore, the possibility of flow stripping occurring in the first inclined surface 61A can also be reduced.

[0056] More specifically, because the nozzle width at the blade tip is greater than that at the middle section 50M, the amount of power acquired from the combustion gas by the last-stage moving blade row 50 is reduced at the blade tip. On the other hand, because the outflow angle at the middle section 50M is larger (i.e., the nozzle width is smaller), the amount of power acquired from the combustion gas is increased at the middle section 50M. Here, as... Figure 7As shown by the double-dotted line, conventionally, the total exhaust pressure at the diffuser 60 inlet is approximately constant from the tip side of the last-stage moving blade row to the hub side, and peeling easily occurs during exhaust flow near the wall of the outer cylinder 61 or inner cylinder 62. As a result, there is a problem of reduced static pressure recovery in the diffuser 60. In contrast, in this embodiment, as... Figure 7 As shown by the solid line, the total exhaust pressure at the inlet of the diffuser 60 on the blade tip side is higher than that of the middle portion 50M of the moving blade body 50H. Therefore, the angle θ2 formed by the first inclined surface 61A with respect to the axis Ac can be further increased compared to the past. As a result, further stripping is less likely to occur in the exhaust flow near the walls of the outer cylinder 61 and the inner cylinder 62. Consequently, the static pressure recovery in the diffuser 60 can be further increased.

[0057] Furthermore, according to the above structure, in the last-stage moving blade row 50, the nozzle width A1 at the radially inner end (hub side) is greater than the nozzle width A2 at the middle part 50M. As a result, the total pressure on the hub side, in addition to the blade tip side, also increases. Therefore, the possibility of flow stripping occurring on the outer circumferential surface 62A of the inner cylinder 62 can be reduced. Consequently, the static pressure recovery in the diffuser 60 can be further increased, and the performance of the gas turbine 10 can be further improved.

[0058] Furthermore, in the above structure, the area ratio of the diffuser's flow path cross-sectional area (cross-sectional area S1) in the upstream end edge (leading edge 63A) of the support 63 to the flow path cross-sectional area (cross-sectional area S2) at the inlet of the diffuser 60 is set to be 1.28 or more and 1.37 or less, which is larger than that of conventional gas turbines. As a result, the exhaust velocity can be further reduced in the region upstream of the support 63.

[0059] Furthermore, in the last-stage moving blade row 50, when viewed circumferentially, the blade tip side end face 50C extends radially outward from the upstream side towards the downstream side. The angle θ1 formed by the blade tip side end face 50C relative to the axis Ac is set to be greater than the angle θ2 formed by the first inclined surface 61A relative to the axis Ac. The flow component of the exhaust gas flowing radially outward along the blade tip side end face 50C is guided by the first inclined surface 61A located on the downstream side. Because the angle θ1 formed by the blade tip side end face 50C relative to the axis Ac is greater than the angle θ2 formed by the first inclined surface 61A relative to the axis Ac, the aforementioned flow component is pressed tightly against the first inclined surface 61A from the radial inward side. As a result, the generation of flow stripping in the first inclined surface 61A can be further suppressed. Thus, the generation of vortices in the first inclined surface 61A can be avoided.

[0060] Furthermore, in the above structure, the difference between the angle θ1 formed by the blade tip end face 50C relative to the axis Ac and the angle θ2 formed by the first inclined surface 61A relative to the axis Ac is 0° or more and 5° or less. Here, as described above, since the nozzle width in the blade tip end face 50C is greater than the nozzle width in the middle portion 50M, the total exhaust pressure in the blade tip side can be maintained more effectively. As a result, the difference between the aforementioned angles θ1 and θ2 can be made smaller than before. In other words, as long as angle θ1 is greater than angle θ2, the angle θ2 formed by the first inclined surface 61A relative to the axis Ac can be expanded to the maximum allowable limit. As a result, flow stripping in the first inclined surface 61A can be further suppressed.

[0061] Furthermore, in the above structure, the support column 63 is configured to span the first inclined surface 61A and the second inclined surface 61B in the direction of the axis Ac. Therefore, before contacting the support column 63, the exhaust flow is sufficiently decelerated in the region on the first inclined surface 61A side. As a result, pressure loss generated by the support column 63 can be further reduced.

[0062] (Other implementation methods)

[0063] The embodiments of the present invention have been described above. Furthermore, various changes or modifications can be made to the above structures without departing from the spirit of the invention. For example, in the above embodiments, an example was described where the support column 63 is arranged across the first inclined surface 61A and the second inclined surface 61B of the outer cylinder 61. However, the arrangement of the support column 63 is not limited to the above. For example, a structure in which the support column 63 is arranged only on the second inclined surface 61B can also be used. According to this structure, by flowing along the first inclined surface 61A, the exhaust velocity can be further reduced in a region upstream of the support column 63. As a result, the pressure loss generated by the support column 63 can be further reduced.

[0064] Furthermore, in the above embodiment, only the nozzle width and discharge angle of the last-stage moving blade row 50 have been described. However, it is also possible to construct a last-stage stationary blade row disposed upstream of the last-stage moving blade row 50 by having the same nozzle width and discharge angle as the last-stage moving blade row 50.

[0065] <Postscript>

[0066] The turbine 40 and gas turbine 10 described in each embodiment are, for example, as follows.

[0067] (1) The turbine 40 according to the first embodiment includes: a turbine rotor 41 extending along an axis Ac and rotatable about the axis Ac; a turbine housing 45 covering the turbine rotor 41 from the outer peripheral side; a plurality of turbine moving blade rows 43 having a plurality of turbine moving blades arranged circumferentially along the axis Ac on the outer peripheral surface of the turbine rotor 41 and arranged along the axis Ac direction; a plurality of turbine stationary blade rows 46 having a plurality of turbine stationary blades arranged circumferentially on the inner peripheral surface of the turbine housing 45 in an adjacent manner to the turbine moving blades on one side in the axis Ac direction and arranged along the axis Ac direction; and a diffuser 60 disposed on the other side of the last stage moving blade row 43A in the axis Ac direction of the plurality of turbine moving blade rows 43, and forming an exhaust flow in which exhaust gas flows from one side to the other side in the axis Ac direction. The diffuser 60 has: an inner cylinder 62 extending along the axis Ac; an outer cylinder 61 covering the inner cylinder 62 from the outer periphery and forming the exhaust flow path E between the outer cylinder 61 and the inner cylinder 62; and a plurality of supports 63 disposed at the midpoint of the exhaust flow path E, connecting the inner cylinder 62 and the outer cylinder radially and arranged circumferentially. The outer cylinder 61 has a first inclined surface 61A extending radially inward from the axis Ac as the inlet of the exhaust flow path E on one side faces the other side. The first inclined surface 61A forms an angle of 16° or more and 24° or less relative to the axis Ac in a cross-section including the axis Ac. In the turbine blades of the last stage moving blade row 43A, the nozzle width A1 at the radially outer end relative to the axis Ac is set to be greater than the nozzle width A2 of the radially middle portion 50M.

[0068] According to the above structure, the angle between the first inclined surface 61A and the axis Ac is set to be 16° or more and 24° or less, which is larger than that of a conventional diffuser. This allows for a further reduction in the exhaust velocity flowing through the exhaust path E in the region closer to the axis Ac than the support 63. As a result, pressure loss generated by the support 63 can be further suppressed and reduced. Furthermore, in the turbine blades of the last-stage moving blade row 43A, the nozzle width A1 at the radially outer end is greater than the nozzle width A2 at the middle portion 50M. This results in a higher total pressure at the radially outer end. Therefore, the possibility of flow stripping in the first inclined surface 61A can also be reduced.

[0069] (2) In the turbine 40 involved in the second method, in the turbine blades of the last stage moving blade row 43A, the nozzle width A1 of the radially inner end is set to be greater than the nozzle width A2 of the radially upper middle part 50M.

[0070] According to the above structure, in the turbine blades of the last-stage moving blade row 43A, the nozzle width A1 at the radially inner end is greater than the nozzle width A2 at the middle part 50M. As a result, the total pressure at the radially inner end is higher. Therefore, the possibility of flow stripping occurring on the outer circumferential surface 62A of the inner cylinder 62 can be reduced.

[0071] (3) In the turbine 40 involved in the third method, the ratio of the flow path cross-sectional area of ​​the diffuser 60 on the end edge of the support 63 in the direction of the axis Ac to the flow path cross-sectional area of ​​the diffuser 60 at the inlet on the direction of the axis Ac is 1.28 or more and 1.37 or less.

[0072] According to the above structure, the exhaust flow rate can be further reduced in the region on one side of the axis Ac direction of the support column 63.

[0073] (4) In the turbine 40 involved in the fourth method, when viewed from the circumferential direction, the radially outer tip end face 50C extends radially outward from one side to the other in the direction of the axis Ac, and the angle θ1 of the tip end face 50C relative to the axis Ac is set to be greater than the angle θ2 of the first inclined surface 61A relative to the axis Ac.

[0074] According to the above structure, the flow component along the tip-side end face 50C towards the radially outward direction is guided by the first inclined surface 61A on the other side of the axis Ac. Since the angle θ1 formed by the tip-side end face 50C with respect to the axis Ac is greater than the angle θ2 formed by the first inclined surface 61A with respect to the axis Ac, the flow component is pressed tightly against the first inclined surface 61A. As a result, the stripping of the flow in the first inclined surface 61A can be further suppressed.

[0075] (5) In the turbine 40 involved in the fifth method, the difference between the angle θ1 formed by the blade tip side end face 50C relative to the axis Ac and the angle θ2 formed by the first inclined surface 61A relative to the axis Ac is more than 0° and less than 5°.

[0076] According to the above structure, the stripping of the flow in the first inclined surface 61A can be further suppressed.

[0077] (6) In the turbine 40 of the sixth embodiment, the outer cylinder 61 further has a second inclined surface 61B, which is connected to the other side of the first inclined surface 61A in the direction of the axis Ac, and extends radially from the inner side to the outer side as it moves from one side of the axis Ac to the other side. In a cross-section including the axis Ac, the angle between the second inclined surface 61B and the axis Ac is smaller than the angle θ2 between the first inclined surface 61A and the axis Ac. The support 63 is configured to span the first inclined surface 61A and the second inclined surface 61B in the direction of the axis Ac.

[0078] According to the above structure, the support column 63 is configured to span the first inclined surface 61A and the second inclined surface 61B. Therefore, before contacting the support column 63, the flow of exhaust gas is sufficiently slowed in the region on the first inclined surface 61A side. As a result, pressure loss generated by the support column 63 can be further reduced.

[0079] (7) The gas turbine 10 involved in the seventh method includes: a compressor 20 that generates high-pressure air by generating compressed air A; a burner 30 that mixes fuel with the high-pressure air to generate combustion gas G; and a turbine 40 that is driven by the combustion gas G.

[0080] Based on the above structure, a gas turbine 10 can be provided that further improves performance by further reducing pressure loss in the diffuser 60.

[0081] Industrial availability

[0082] According to the present invention, a turbine and gas turbine can be provided that further improves performance by reducing pressure loss.

[0083] Symbol Explanation

[0084] 10-Gas turbine, 11-Gas turbine rotor, 15-Gas turbine casing, 16-Intermediate chamber, 20-Compressor, 21-Compressor rotor, 22-Compressor rotating shaft, 23-Compressor moving blade row, 25-Compressor casing, 26-Compressor stationary blade row, 30-Burner, 40-Turbine, 41-Turbine rotor, 42-Turbine rotating shaft, 43-Turbine moving blade row, 43A-Last stage moving blade row, 45-Turbine casing, 45A-Inner circumferential surface, 46-Turbine stationary blade row, 50-Last stage moving blade row, 50A-Leading edge, 50B-Leftward edge, 50C- 50D-Booth side end face, 50H-Moving blade body, 50M-Intermediate part, 50N-Negative pressure surface, 50P-Positive pressure surface, 60-Diffuser, 61-Outer cylinder, 61A-First inclined surface, 61B-Second inclined surface, 62-Inner cylinder, 62A-Outer peripheral surface, 63-Support column, 63A-Support column leading edge, 63B-Support column trailing edge, 64-Manhole, 70-Disc, 70A-Outer peripheral surface, 80-Bearing device, Ac-Axis axis, A-Air, E-Exhaust flow path, F-Fuel, G-Combustion gas, θ1, θ2-Angles, θ3, θ4-Outflow angles.

Claims

1. A gas turbine comprising: Turbines, including: A turbine rotor that extends along an axis and is capable of rotating about the axis; And a plurality of turbine stationary blade rows and a plurality of turbine moving blade rows arranged alternately one by one in the axial direction of the turbine rotor; and The diffuser is disposed downstream of the last stage blade row, located on the most downstream side, in the direction of combustion gas flow along the axis of the plurality of turbine blade rows, and forms an exhaust flow path for exhaust gas to flow from the inlet to the outlet. The diffuser has: Inner cylinder; An outer cylinder covers the inner cylinder from its outer periphery and forms the exhaust flow path between the outer cylinder and the inner cylinder; and Multiple support pillars are positioned midway along the exhaust flow path, arranged circumferentially along the axis, and radially connect the inner cylinder and the outer cylinder. The outer cylinder has a first inclined surface that extends radially inward from the inlet side centered on the axis toward the outlet side in the axial direction of the exhaust flow path. The first inclined plane, in a cross-sectional view including the axis, forms an angle of 16° or more and 24° or less relative to the axis. In the plurality of turbine blades constituting the final stage blade array, the nozzle width at the radially outer end relative to the axis is set to be greater than the nozzle width at the radially middle portion. In each of the turbine blades constituting the final stage blade row, the radially outer tip end face extends radially outward in the direction of combustion gas flow along the axis from the upstream side to the downstream side, and the angle formed by the tip end face relative to the axis is set to be greater than the angle formed by the first inclined surface relative to the axis.

2. The gas turbine according to claim 1, wherein, In the plurality of turbine blades constituting the final stage blade array, the nozzle width at the radially inner end is set to be greater than the nozzle width at the radially middle portion.

3. The gas turbine according to claim 1 or 2, wherein, The ratio of the flow path cross-sectional area S2 of the diffuser located on the front edge of the inlet side of the support in the axial direction to the flow path cross-sectional area S1 at the inlet of the diffuser, S2 / S1, is greater than 1.28 and less than 1.

37.

4. The gas turbine according to claim 1 or 2, wherein, The difference between the angle formed by the blade tip side face relative to the axis and the angle formed by the first inclined surface relative to the axis is greater than 0° and less than 5°.

5. The gas turbine according to claim 1 or 2, wherein, The outer cylinder also has a second inclined surface connected to the axial direction outlet side of the first inclined surface and extending radially from the inner side to the outer side from the axial direction inlet side toward the axial direction outlet side. In a cross-section including the axis, the angle between the second inclined surface and the axis is smaller than the angle between the first inclined surface and the axis. The support is configured to span the first inclined surface and the second inclined surface in the axial direction.

6. The gas turbine according to claim 1 or 2, wherein, In each of the turbine blades constituting the final stage moving blade row, the angle between the tip end face and the axis is more than 20° and less than 25°.

7. The gas turbine according to claim 1 or 2, comprising: A compressor generates high-pressure air by compressing air; and A burner that mixes fuel with the high-pressure air to generate combustion gases.

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