JP2011163184A - Outlet blade cascade of gas turbine compressor - Google Patents

Abstract

<P>PROBLEM TO BE SOLVED: To provide an outlet blade cascade of a gas turbine compressor capable of suppressing the wear of a fitting portion of a stator blade cascade in the compressor. <P>SOLUTION: The outlet blade cascade 3 of the gas turbine compressor 57 includes an outer ring 5 inserted in a turbine circumferential direction and fitted in the inner peripheral part of a turbine casing 1, an inner ring 7 inserted in the turbine circumferential direction and fitted in a recessed groove 11 of an inner barrel 4 enclosing a turbine rotating shaft, a plurality of stator blades 6 connected at a tip side to the outer ring 5 and at a root side to the inner ring 7, and a spacer 14 provided between the turbine casing 1 and the outer ring 5 and formed of a material higher in the coefficient of thermal expansion than the outer ring 5. <P>COPYRIGHT: (C)2011,JPO&INPIT

Description

  The present invention relates to an outlet cascade of a gas turbine compressor.

  In a gas turbine facility, a stationary blade cascade of a compressor is formed in an annular shape by connecting a plurality of sector segments in a circumferential direction, and is arranged in a plurality of stages in the axial direction between a turbine casing and an inner barrel. The fan-shaped segment of each stationary blade cascade has a configuration in which the leading end side of the stationary blade is fixed to the outer ring and the root side is fixed to the inner ring (see Patent Document 1, etc.).

JP 2005-194903 A

  In the case of a compressor of a gas turbine facility, a plurality of stages of stationary blade cascades (hereinafter referred to as outlet blade rows) may be provided at an outlet portion in order to rectify compressed air and introduce the compressed air into a combustion apparatus. In such a fan-shaped segment of the stationary vane at the outlet of the compressor, the hook provided on the outer ring is fitted on the hook groove of the turbine casing on the outer peripheral side, and the inner ring is fitted on the inner barrel on the inner peripheral side.

  However, each fitting portion between the fan-shaped segment and the turbine casing or the inner barrel requires a certain assembly gap in consideration of assemblability. For this reason, the rotational force in the circumferential direction, the pressing force in the axial direction, and the external force due to the rotational vibration of the rotating shaft are received by the fluid force of the compressed air during operation, and the fan-shaped segment fitting part moves in the gap. The outer ring and the inner ring wear due to friction with the turbine casing and the inner barrel. Further, when the wear progresses, the movement of the fan-shaped segment increases due to an increase in the gap size, which causes further wear.

  An object of the present invention is to provide an outlet cascade of a gas turbine compressor capable of suppressing wear of a fitting portion of a stationary blade cascade of a compressor.

  In order to achieve the above object, a first invention is an outlet cascade of a gas turbine compressor, an outer ring inserted and fitted in an inner peripheral portion of a turbine casing in a turbine circumferential direction, and a turbine rotating shaft An inner ring that is inserted and fitted in a recess in the inner barrel surrounding the inner barrel, a plurality of stationary blades having a tip side connected to the outer ring and a root side connected to the inner ring, the turbine casing, and the outer ring And a spacer formed between a ring and a material having a higher coefficient of thermal expansion than the outer ring.

  A second invention is an outlet blade row of a gas turbine compressor, and includes an outer ring that is inserted and fitted in an inner peripheral portion of a turbine casing in a circumferential direction of the turbine, and a concave portion of an inner barrel that surrounds the turbine rotating shaft. An inner ring that is inserted and fitted in the turbine circumferential direction, a plurality of stationary blades having a tip side connected to the outer ring and a root side connected to the inner ring, and provided between the inner barrel and the inner ring, And a spacer made of a material having a higher coefficient of thermal expansion than the inner ring.

  According to a third invention, in the first or second invention, the spacer is formed in a ring shape.

  According to a fourth invention, in the first or second invention, the spacer is partially disposed in a circumferential direction of the turbine.

  A fifth invention is characterized in that, in the first or second invention, the outer ring is divided into a main body part connected to the stationary blade and a fitting part fitted to the turbine casing. To do.

  A sixth invention is characterized in that, in the first or second invention, the inner ring is divided into a main body part connected to the stationary blade and a fitting part fitted to the inner barrel. To do.

  According to a seventh invention, in the first or second invention, the inner ring has a slit extending in the circumferential direction and having a depth in the turbine radial direction.

  According to an eighth aspect of the present invention, in the first or second aspect of the invention, a surface hardening process is performed on each of the facing portions of the outer ring and the inner ring facing the turbine casing and the inner barrel.

  A ninth invention is characterized in that, in the first or second invention, the outer ring has a slit extending in the turbine axial direction and having a depth in the turbine radial direction.

  According to a tenth invention, in the first or second invention, an elastic member is disposed between parts adjacent to each other in the circumferential direction of the inner ring.

  ADVANTAGE OF THE INVENTION According to this invention, abrasion of the fitting part of the stationary blade cascade of a compressor can be suppressed.

It is a fragmentary sectional side view of the example of 1 composition of the gas turbine to which the outlet cascade of the gas turbine compressor of the present invention is applied. (A) It is sectional drawing by the side of an expansion of the exit part of a gas turbine compressor, and (b) Sectional drawing by the II-II arrow in Drawing 2 (a). (A) The figure which looked at the basic structure of the fan-shaped segment of the outlet blade row | line | column of the gas turbine compressor which concerns on 1st Embodiment of this invention from the turbine axial direction, (b) From the arrow III direction in Fig.3 (a) FIG. (A) The figure which looked at the fan-shaped segment of the exit blade row | line | column of the gas turbine compressor which concerns on 1st Embodiment of this invention from the turbine axial direction, (b) The figure seen from the arrow IV direction in Fig.4 (a) It is. (A) The figure which looked at the fan-shaped segment of the exit blade row | line | column of the gas turbine compressor which concerns on 2nd Embodiment of this invention from the turbine axial direction, (b) The figure seen from the arrow V direction in Fig.5 (a) It is. (A) The figure which looked at the fan-shaped segment of the exit blade row | line | column of the gas turbine compressor which concerns on 3rd Embodiment of this invention from the turbine axial direction, (b) The figure seen from the arrow VI direction in Fig.6 (a) It is. (A) The figure which looked at the fan-shaped segment of the outlet blade row | line | column of the gas turbine compressor which concerns on 4th Embodiment of this invention from the turbine axial direction, (b) The figure seen from the arrow VII direction in Fig.7 (a) It is. (A) The figure which looked at the fan-shaped segment of the exit blade row | line | column of the gas turbine compressor which concerns on 5th Embodiment of this invention from the turbine axial direction, (b) The figure seen from the arrow VIII direction in Fig.8 (a) It is. (A) The figure which looked at the fan-shaped segment of the exit blade row | line | column of the gas turbine compressor which concerns on 6th Embodiment of this invention from the turbine axial direction, (b) The figure seen from the arrow IX direction in Fig.9 (a) It is. (A) The figure which looked at the fan-shaped segment of the exit blade row | line | column of the gas turbine compressor which concerns on 7th Embodiment of this invention from the turbine axial direction, (b) The figure seen from the arrow X direction in Fig.10 (a) It is. (A) The figure which looked at the fan-shaped segment of the outlet cascade of the gas turbine compressor which concerns on 8th Embodiment of this invention from the turbine axial direction, and (b) of the gas turbine compressor which concerns on 8th Embodiment of this invention. It is a perspective view of the elastic member with which the exit blade row was equipped. (A) The figure which looked at the fan-shaped segment of the exit blade row | line | column of the gas turbine compressor which concerns on 9th Embodiment of this invention from the turbine axial direction, (b) The figure seen from the arrow XII direction in Fig.12 (a) It is.

  Embodiments of the present invention will be described below with reference to the drawings.

<First Embodiment>
FIG. 1 is a partial cross-sectional side view of a configuration example of a gas turbine to which an outlet cascade of a gas turbine compressor of the present invention is applied.

  The gas turbine shown in FIG. 1 includes an axial compressor 57 that compresses air 2a, a combustion device 54 that combusts compressed air 2b from the axial compressor 61 together with fuel, and a combustion gas from the combustion device 54. And a turbine 52 driven by The axial flow type compressor 57 and the rotating shaft (rotor) 55 of the turbine 52 are connected coaxially and accommodated in the turbine casing 1.

  The turbine 52 includes a rotating shaft 55, a moving blade 51 provided on the inner peripheral portion of the rotating shaft 55, and a stationary blade 50 provided on the inner peripheral portion of the turbine casing 1. A plurality of moving blades 51 and stationary blades 50 are provided in the turbine rotating direction to form an annular blade row, and the blade rows of the moving blades 51 and the stationary blades 50 are alternately arranged in the axial direction.

  The air compressor 53 includes a rotary shaft 55, an air compressor rotor blade 56 provided on the outer peripheral portion of the rotary shaft 55, and an air compressor stationary blade 53 provided on the inner peripheral portion of the turbine casing 1. A plurality of air compressor blades 56 and air compressor stationary blades 53 are provided in the turbine rotation direction to form an annular blade row, and the blade rows of the air compressor blades 56 and the air compressor stationary blades 53 are alternately arranged in the axial direction. is set up.

  When the rotating shaft 55 rotates, the air 2a is sucked into the axial flow compressor 57 and is compressed while turning. The high-temperature and high-pressure compressed air 2 b generated by the axial-flow compressor 57 is rectified into an axial flow by the compressor outlet blade row 3 provided at the outlet of the axial-flow compressor 57 and introduced into the combustion device 54. Is done. The compressed air 2b (or compressed air) from the axial flow type compressor 57 is extracted as cooling air for the high temperature member at various locations in the gas turbine.

  The combustion device 54 burns fuel together with the compressed air 2 b to generate a combustion gas H of about 1300 ° C. and supplies it to the turbine 52. The combustion gas H supplied to the turbine 52 is injected into the moving blade 51 through the stationary blade 50 in each stage, and thereby the turbine 52 is driven. The combustion gas H that has driven the turbine 52 is discharged to the outside as an exhaust gas of about 650 ° C. through an exhaust diffuser (not shown) that constitutes the flue. Although not particularly illustrated, a rotor of a load device such as a generator is connected to the rotary shaft 55, and the load device is driven by the rotational power of the rotary shaft 55 obtained as described above. Further, the entire gas turbine is covered with a gas turbine enclosure (not shown) in order to cut off the influence of heat on the outside and to prevent sound.

  2A is an enlarged side sectional view of the outlet portion of the axial compressor 57, and FIG. 2B is a sectional view taken along the line II-II in FIG. 2A.

  As described above, the outlet of the axial flow type compressor 57 has an annular compressor outlet for rectifying the swirling flow of the compressed air 2 parallel to the axial direction and introducing it into the combustion device 54. A cascade 3 is provided. The compressor outlet blade row 3 is inserted in the turbine circumferential direction into the outer ring 5 that is inserted and fitted into the inner circumferential portion of the turbine casing 1 in the turbine circumferential direction, and the inner barrel 4 that surrounds the rotary shaft 55. And a plurality of stationary blades (rectifying blades) 6 whose tip side is connected to the inner ring 7 and whose root side is connected to the inner ring 7. In the present embodiment, the compressor outlet blade row 3 is formed in an annular shape by arranging a plurality of fan-shaped segments in the circumferential direction. That is, both the outer ring 5 and the inner ring 7 are divided into a plurality of parts in the circumferential direction, and the fan-shaped segment is formed by connecting the parts facing the outer ring 5 and the inner ring 7 in the radial direction with a plurality of stationary blades 6. It is a configuration. The compressor outlet blade row 3 is arranged in a plurality of stages in the axial direction so that the stationary blade 6 faces the flow path of the compressed air 2 in the downstream of the compressor rotor blade 56 at the final stage of the axial flow compressor 57. The compressor outlet blade rows 3 in each stage have the same configuration.

  FIG. 3A is a view of the basic structure of the fan-shaped segment of the outlet blade row of this embodiment as seen from the turbine axis direction, and FIG. 3B is a view of this as seen from the direction of arrow III in FIG. It is.

  As shown in FIGS. 3 (a) and 3 (b), a blade dovetail portion 9 is provided at the tip (outer peripheral side) of the stationary blade 6, and the blade dovetail portion 9 is axially (assembled). It is fitted into dovetail grooves provided at regular intervals in the circumferential direction on the inner circumferential part of the outer ring 5 from the axial direction in the state, and after fitting, it is fixed so that it does not come out of the outer ring 5 by caulking. The outer ring 5 has an outer ring hook 8 on the outer periphery, and the outer ring hook 8 is fitted into the turbine casing 1 by being inserted into a hook groove 13 provided in the turbine casing 1 from the circumferential direction. ing. The root portion (inner peripheral side) of the stationary blade 6 is fixed to the outer peripheral portion of the inner ring 7 by welding. The inner ring 7 is fitted in an annular concave groove 11 provided on the outer peripheral portion of the inner barrel 4. The inner ring 7 is provided with a certain gap 12 between the divided parts adjacent in the circumferential direction in order to attenuate the fluid force of the compressed air 2 b applied to the stationary blade 6. The stationary blade 6 can be displaced in the circumferential direction within the range of the gap 12. The outer ring 5 and the inner ring 7 are substantially flush with the hook groove 13 and the concave groove 11, respectively, so as not to disturb the flow of the compressed air 2b.

  Here, the hook groove 13 of the turbine casing 1 and the concave groove 11 of the inner barrel 4 are formed larger than the outer ring 5 and the inner ring 7, respectively, in consideration of assembly. Between the hook groove 13 of the turbine casing 1 and the outer ring hook 8, for example, a gap of about 0.1 mm is interposed between the concave groove 11 of the inner barrel 4 and the inner ring 7, for example, about 2 mm. The compressor outlet blade row 3 has a slight backlash with respect to the turbine casing 1 and the inner barrel 4. Therefore, the fan-shaped segment is axially, radially, or circumferentially affected by the fluid force of the compressed air 2b during operation, and receives the rotational force in the circumferential direction, the pressing force in the axial direction, or the external force due to the rotational vibration of the rotary shaft 55. As a result, the frictional force is generated between the turbine casing 1 and the inner barrel 4 as shown in FIG. 3B, and the axial direction of the outer ring hook 8 and the outer ring 5 is directed. Wear A, B, C occurs on the surfaces 5a, 5b, the surfaces 7a, 7b facing the axial direction of the inner ring 7, and the like. Further, when the wear AC progresses, the wear further proceeds due to an increase in the movement amount of the fan-shaped segments. In particular, the hook groove 13 of the turbine casing 1 is formed with a machining relief groove 10 for machining the hook groove 13 on the outer peripheral side. Therefore, when the part A of the outer ring hook 8 advances, the radial direction of the fan-shaped segment is increased. As a result, the compressor outlet blade 3 is easily displaced in the radial direction, which causes other wear B and C to accelerate.

  Further, since the ambient temperature on the outer peripheral side of the turbine casing 1 in the gas turbine enclosure (not shown) is about 80 ° C., the compressed air 2b of about 450 ° C. flows through the axial flow compressor 57. Due to the temperature difference, the cross section of the turbine casing 1 in the axial flow compressor 57 portion tends to be thermally deformed in an elliptical shape (see the two-dot chain line in FIG. 2). As a result, an external force due to thermal deformation of the turbine casing 1 is applied to the outer ring 5, and a strong contact force is generated between the outer ring hook 8 and the hook groove 13 of the turbine casing 1. This can also be a factor for promoting wear of the outer ring hook 8.

  When wear AC occurs in this way, it is necessary to restore the shapes of the outer ring 5 and the inner ring 7 by welding repair at every maintenance. This welding repair is not only a troublesome work requiring a lot of labor and time, but if the amount of thermal deformation becomes significant due to the amount of welding, the outer ring 5 and the inner ring 7 cannot be restored to their original form, and the economic efficiency of the gas turbine deteriorates. Let

  When the outer ring hook 8 is worn out, the fan-shaped segment of the compressor outlet blade row 3 is disengaged from the turbine casing 1, and the fan-shaped segment cannot be held at the normal position. As a result, if the compressed air 2b is not properly rectified, the predetermined properties of the compressed air 2b cannot be obtained, and the performance of the gas turbine is also deteriorated. In addition, if the compressed air 2b is introduced into the combustion device 54 without being properly rectified, there is a risk of abnormal combustion, and the reliability of the gas turbine is reduced.

  As described above, the wear of the outer ring 5 and the inner ring 7 of the compressor outlet cascade 3 is a serious problem in terms of plant reliability and economy.

  Therefore, in the present embodiment, the wear generated in the compressor outlet blade row 3 is suppressed by the measures described below, the deterioration of the function of the gas turbine parts and the occurrence of welding repair work are suppressed, the reliability of the gas turbine and The economy is being improved.

  FIG. 4A is a view of the fan-shaped segment of the outlet blade row of the gas turbine compressor according to the first embodiment of the present invention as viewed from the turbine axial direction, and FIG. 4B is the view in FIG. 4A. It is the figure seen from the arrow IV direction. The same parts as those in the previous drawings are denoted by the same reference numerals as those used in the previous drawings, and the description thereof is omitted.

  As shown in FIGS. 4A and 4B, in the present embodiment, the spacer 14 is provided in the above-described machining relief groove 10 for machining the hook groove 13 in the turbine casing 1. The spacer 14 is formed of a material (aluminum or the like) having a higher thermal expansion coefficient than the outer ring 5, and is attached to the outer peripheral portion of the outer ring 5 with a fixing screw 15, and is arranged between the turbine casing 1 and the outer ring 5. It is installed. The spacer 14 may be fixed to the outer ring 5 by other methods such as welding, or may not be fixed to the outer ring 5 but simply accommodated in the machining clearance groove 10. The thickness of the spacer 14 is equivalent to the depth dimension in the turbine radial direction of the machining relief groove 10 taken from the boundary with the hook groove 13, and the dimension in the turbine axial direction is slightly smaller than that of the machining relief groove 10. It is.

  Further, the spacer 14 of the present embodiment is formed in a ring shape. At this time, in the case where the spacer 14 is not fixed to the fan-shaped segment and is incorporated in the turbine casing 1 prior to the fan-shaped segment, the spacer 14 may be an integral ring-shaped member. Although it is good also as a crack structure, when integrating in the turbine casing 1 with a fan-shaped segment, the structure which divides | segments into several parts corresponding to each fan-shaped segment, for example, and fixes to a fan-shaped segment is preferable. Although not shown in particular, the spacer 14 and the outer ring 5 may be provided with a so-called inlay structure so that the positional relationship between them can be reproduced with high accuracy.

  According to the present embodiment, since the compressed air 2b flowing into the compressor outlet blade row 3 is at a high temperature of about 450 ° C., a thickness equivalent to the radial depth of the machining relief groove 10 on the outer peripheral side of the outer ring 5 is achieved. By providing the spacers 14 which are the protrusions, the spacers 14 are thermally expanded in the turbine radial direction so that the fan-shaped segments are pressed to the inner peripheral side. As a result, since the displacement of the sector segment in the turbine radial direction is constrained, it is possible to suppress the progress of wear of the fitting portion of the sector segment with the turbine casing 1 or the inner barrel 4. Further, when the spacer 14 is fixed to the outer ring 5, when the spacer 14 is thermally expanded in the turbine shaft direction and restrained in the machining relief groove 10, the sector segment is also displaced in the turbine shaft direction. It can also be restrained, and the progress of wear of the fan-shaped segment is further suppressed.

  As described above, according to the present embodiment, the spacer 14 that is more thermally expanded than the outer ring 5 is provided on the outer peripheral portion of the outer ring 5, so that the fluid force of the compressed air 2 b and the rotational vibration of the rotary shaft 55 are affected. The wear of the fitting part of the fan-shaped segment resulting from it can be suppressed. As described above, since the wear of the fitting portion of the fan-shaped segment can be suppressed, it is possible to suppress the occurrence of welding repair work during the periodic inspection or the amount of repair welding, and to improve the economic efficiency of the gas turbine. Moreover, since wear of the outer ring hook 8 can be suppressed, displacement of the compressor outlet blade row 3 can be suppressed. Thereby, since the fall of the rectification effect of compressed air 2b can be controlled, the fall of the reliability of a gas turbine is also controlled.

  Further, since the spacer 14 is accommodated and disposed in the machining clearance groove 10 formed along with the hook groove 13 of the turbine casing 1 which is the basic structure, it is also advantageous in that it can be easily applied to an existing plant. is there.

Second Embodiment
FIG. 5A is a view of the fan-shaped segment of the outlet blade row of the gas turbine compressor according to the second embodiment of the present invention as seen from the turbine axial direction, and FIG. 5B is the view in FIG. 5A. It is the figure seen from the arrow V direction. The same parts as those in the previous drawings are denoted by the same reference numerals as those used in the previous drawings, and the description thereof is omitted.

  The present embodiment is different from the first embodiment in that the spacer 14A is not connected in a ring shape and is partially provided on the outer peripheral portion of the outer ring 5 as shown in the drawing. Although the circumferential dimension and spacing of the spacer 14A are not particularly limited, in FIG. 5A, the circumferential dimension of the spacer 14A is about 1 pitch in the circumferential direction of the stationary blade 6, and the spacing is 3 pitch in the circumferential direction of the stationary blade 6. The case of the degree is illustrated. Other configurations such as the cross-sectional shape of the spacer 14A are the same as in the first embodiment.

  According to the present embodiment, substantially the same effect as that of the first embodiment can be obtained, and the spacer 14A is partially provided on the outer peripheral portion of the outer ring 1, so that compared to the first embodiment, The material cost of the spacer 14A can be reduced, and the weight of the spacer 14A can be reduced to reduce the labor for attaching and detaching.

<Third Embodiment>
FIG. 6A is a view of the fan-shaped segment of the outlet blade row of the gas turbine compressor according to the third embodiment of the present invention as viewed from the turbine axial direction, and FIG. 6B is the view in FIG. 6A. It is the figure seen from the arrow VI direction. The same parts as those in the previous drawings are denoted by the same reference numerals as those used in the previous drawings, and the description thereof is omitted.

  This embodiment is different from the above-described embodiments in that the outer ring 5A is divided into a main body portion 5Aa connected to the stationary blade 6 by welding, for example, and a fitting portion 5Ab fitted to the turbine casing 1. It is a point.

  The fitting portion 5Ab is attached to the main body portion 5Aa by a method that can be easily detached like the fixing screw 15. The position of the boundary between the main body portion 5Aa and the fitting portion 5Ab is not particularly limited, but ensures a necessary and sufficient strength. In the present embodiment, both the main body 5Aa and the fitting portion 5Ab are divided so that the cross section is rectangular. In addition, a so-called inlay structure may be provided in the main body 5Aa and the fitting portion 5Ab so that the positional relationship between the two can be reproduced with high accuracy. The spacer 14 (which may be the spacer 14A of the second embodiment) is located on the outer peripheral portion of the fitting portion 5Ab of the outer ring 5A, and is attached to the main body portion 5Aa together with the fitting portion 5Ab. Other configurations are the same as those of the first embodiment, but the outer ring 5 may be divided into a main body portion and a fitting portion in the second embodiment.

  According to the present embodiment, substantially the same effects as those of the above-described embodiment can be obtained, and even if the fitting portion 5Ab is worn due to the split structure of the fitting portion 5Ab of the outer ring 5A, the wear is caused. It is not necessary to repair the portion by welding, and if only the fitting portion 5Ab is replaced, the original shape can be restored. Therefore, the labor for restoring the original shape can be further reduced, and the economic efficiency of the gas turbine can be improved. Further, since there is no thermal deformation as in the case of welding repair for the wear-thinned portion, the restoration accuracy is high and the reliability of the gas turbine is improved.

<Fourth embodiment>
FIG. 7A is a view of the fan-shaped segment of the outlet blade row of the gas turbine compressor according to the fourth embodiment of the present invention as seen from the turbine axial direction, and FIG. 7B is the view in FIG. 7A. It is the figure seen from the arrow VII direction. The same parts as those in the previous drawings are denoted by the same reference numerals as those used in the previous drawings, and the description thereof is omitted.

  This embodiment is different from the above-described embodiments in that the inner ring 7A is connected to the stator blade 6 by welding or the like, and the fitting portion 18 that fits into the groove 11 of the inner barrel 4. It is a point that is divided.

  The fitting portion 18 is attached to the main body portion 17 by a method that can be easily detached and attached like the fixing screw 25. The position of the boundary between the main body portion 17 and the fitting portion 18 is not particularly limited, but ensures a necessary and sufficient strength. In the present embodiment, the main body portion 17 has a rectangular cross-sectional shape, whereas the fitting portion 18 has a U-shape so as to cover both the surfaces of the main body portion 17 in the turbine axial direction and the inner peripheral surface. Is formed. In addition, a so-called spigot structure may be provided in the main body portion 17 and the fitting portion 18 so that the positional relationship between them can be reproduced with high accuracy. Other configurations are the same as those of the third embodiment, but the inner ring 7 may be divided into a main body portion and a fitting portion in the first or second embodiment.

  Also in the present embodiment, the spacer 14 (or 14A) provided on the outer peripheral side of the outer ring 5 can suppress wear of the fitting portion of the fan-shaped segment as in the above-described embodiments. In addition, since the fitting portion 18 of the inner ring 7A can be attached to and detached from the body portion 17, even if wear occurs in the fitting portion 18, it is not necessary to repair the worn portion by welding. If only the portion 18 is replaced, the original shape can be restored, so that the labor for restoring the original shape can be further reduced and the economic efficiency of the gas turbine can be improved. Further, since there is no thermal deformation as in the case of welding repair for the wear-thinned portion, the restoration accuracy is high and the reliability of the gas turbine is improved.

<Fifth Embodiment>
FIG. 8A is a view of the fan-shaped segment of the outlet blade row of the gas turbine compressor according to the fifth embodiment of the present invention as seen from the turbine axial direction, and FIG. 8B is the view in FIG. 8A. It is the figure seen from the arrow VIII direction. The same parts as those in the previous drawings are denoted by the same reference numerals as those used in the previous drawings, and the description thereof is omitted.

  The present embodiment is different from the above-described embodiments in that an annular slit 19 having a depth in the turbine radial direction and extending in the circumferential direction is provided in the inner ring 7B.

  The depth and width (groove width) of the slit 19 are not particularly limited, but when the surface of the inner ring 7 </ b> B facing the turbine axis direction is strongly pressed against the opposing surface of the concave groove 11, It is necessary to consider so that the part which comprises a surface does not break. Moreover, although the case where the slit 19 was provided in the outer peripheral surface of the inner ring 7B is illustrated in this embodiment, it is good also as a slit which has a depth in a turbine radial direction outer side in an inner peripheral surface. Other configurations are the same as those of the above-described embodiment. Other configurations are the same as those of the third embodiment, but the slit 19 can also be formed in the first, second, or fourth embodiment.

  Also in the present embodiment, the spacer 14 (or 14A) provided on the outer peripheral side of the outer ring 5A can suppress the wear of the fan-shaped segment fitting portion as in the above-described embodiments. In addition, since the slit 19 is provided in the inner ring 7B, the fan-shaped segment is temporarily displaced in the turbine axial direction, and a strong contact force acts between the inner wall of the concave groove 11 of the inner barrel 4 and the inner ring 7. In addition, the contact force is attenuated by bending the portion of the slit 19 that constitutes the wall surface in the turbine axial direction. Therefore, it is possible to more effectively suppress the wear of the inner ring 7B.

<Sixth Embodiment>
FIG. 9A is a view of the fan-shaped segment of the outlet blade row of the gas turbine compressor according to the sixth embodiment of the present invention as viewed from the turbine axial direction, and FIG. 9B is the view in FIG. 9A. It is the figure seen from the arrow IX direction. The same parts as those in the previous drawings are denoted by the same reference numerals as those used in the previous drawings, and the description thereof is omitted.

  This embodiment is different from the above-described embodiments in that a surface hardening process is performed on the facing portions of the outer ring 5 and the inner ring 7B facing the turbine casing 1 and the inner barrel 4.

  That is, in the present embodiment, the outer wall surface of the outer ring 5A that can contact the inner wall surface of the hook groove 13 of the turbine casing 1 and the outer wall surface of the inner ring 7B that can contact the inner wall surface of the recessed groove 11 of the inner barrel 4 are provided. A hardened surface (for example, nitriding) is applied to form a hardened surface layer 20 (shown in bold lines). The part where the hardened surface layer 20 is formed is, for example, a part where the wear AC described above with reference to FIG. Other configurations are the same as those of the fifth embodiment, but the hardened surface layer 20 can also be formed in the first to fourth embodiments.

  Also in the present embodiment, the spacer 14 (or 14A) provided on the outer peripheral side of the outer ring 5A can suppress the wear of the fan-shaped segment fitting portion as in the above-described embodiments. In addition, the surface hardening layer 20 is provided to increase the surface hardness of the contact portion of the outer ring 5A and the inner ring 7B with the turbine casing 1 and the inner barrel 4, thereby increasing the assembly gap between the turbine casing 1 and the inner barrel 4. Without changing, wear of the outer ring 5A and the inner ring 7B can be more effectively suppressed as compared with the case where the surface hardened layer 20 is inner.

<Seventh embodiment>
FIG. 10A is a view of the fan-shaped segment of the outlet blade row of the gas turbine compressor according to the seventh embodiment of the present invention as viewed from the turbine axial direction, and FIG. 10B is the view in FIG. 10A. It is the figure seen from the arrow X direction. The same parts as those in the previous drawings are denoted by the same reference numerals as those used in the previous drawings, and the description thereof is omitted.

  This embodiment is different from the above-described embodiments in that a slit 21 having a depth in the turbine radial direction and extending in the turbine axial direction is provided between adjacent stationary blades 6 of the outer ring 5B.

  The slit 21 is provided on the inner peripheral surface of the outer ring 5B so as to extend outward in the turbine radial direction. In this embodiment, the outer ring 5B is divided into the main body 5Ba and the fitting portion 5Bb. The slits 21 are provided in the main body 5Ba. The depth and width (groove width) of the slit 21 are not particularly limited, but the outer ring 5B (in the case of the present embodiment, the main body 5Ba) does not break when a strong compressive force acts on the outer ring 5B in the circumferential direction. It is necessary to consider. Other configurations are the same as those of the fifth embodiment, but the slits 21 may be formed in the first to fourth or sixth embodiments.

  Also in this embodiment, the spacer 14 (or 14A) provided on the outer peripheral side of the outer ring 5B can suppress the wear of the fan-shaped segment fitting portion as in the above-described embodiments. In addition, since the slit 21 is provided in the outer ring 5B, even if the fan-shaped segments are displaced in the turbine circumferential direction due to thermal deformation of the turbine casing 1 and a strong contact force acts between the fan-shaped segments, Compression in the circumferential direction is allowed and the contact force is attenuated. Therefore, the occurrence of wear on the outer ring 5B can be more effectively suppressed.

<Eighth Embodiment>
FIG. 11A is a view of the fan-shaped segment of the outlet blade row of the gas turbine compressor according to the eighth embodiment of the present invention as viewed from the turbine axial direction, and FIG. 11B is according to the eighth embodiment of the present invention. It is a perspective view of the elastic member with which the exit cascade of the gas turbine compressor was equipped. The same parts as those in the previous drawings are denoted by the same reference numerals as those used in the previous drawings, and the description thereof is omitted.

  The present embodiment is different from the above-described embodiments in that the elastic member 22 is disposed between parts (gap 12) adjacent to each other in the circumferential direction of the inner ring 7B.

  The elastic member 22 is a means for exerting a restoring force in the circumferential direction when compressed between the inner rings 7B, and various springs such as a coil spring can be used as appropriate. The steel plate (dampening plate) bent in is illustrated. The elastic member 22 is fixed to the end surface of one inner ring 7B so as not to hinder its own expansion and contraction operation by a method such as welding, and is not dropped into the compressor flow path. Other configurations are the same as those of the fifth embodiment, but the elastic member 22 may be provided in the first to fourth, sixth, or seventh embodiments.

  Also in the present embodiment, the spacer 14 (or 14A) provided on the outer peripheral side of the outer ring 5A can suppress the wear of the fan-shaped segment fitting portion as in the above-described embodiments. In addition, by providing the elastic member 22 in the gap 12 of the inner ring 7B, the elastic member 22 allows the inner ring 7 to be displaced in the circumferential direction due to the influence of the fluid force of the compressed air 2b. However, the action of strong contact force on the inner ring 7B can be suppressed.

<Ninth Embodiment>
FIG. 12A is a view of the fan-shaped segment of the outlet blade row of the gas turbine compressor according to the ninth embodiment of the present invention as viewed from the turbine axial direction, and FIG. 12B shows this in FIG. 12A. It is the figure seen from the arrow XII direction. The same parts as those in the previous drawings are denoted by the same reference numerals as those used in the previous drawings, and the description thereof is omitted.

  The present embodiment is different from the above-described embodiments in that a spacer 14 is provided between the inner barrel 4 and the inner ring 7.

  In the present embodiment, the spacer 14 is disposed not in the machining escape groove 10 of the turbine casing 1 but in the concave groove 11 of the inner barrel 4, and the spacer 14 is interposed between the inner barrel 4 and the inner ring 7. . The spacer 14 may be fixed to the inner ring 7 with a fixing screw 15 or the like, but is not fixed in this embodiment. The thickness of the spacer 14 is, for example, equal to the depth (turbine radial direction dimension) of the concave groove 16 extending in the circumferential direction on the inner peripheral surface of the inner ring 7, and the turbine axial dimension is larger than that of the concave groove 16. Somewhat small. Although not shown in the present embodiment, it is also conceivable to provide spacers 14 in both the machining clearance groove 10 and the concave groove 11. Other configurations are the same as those of the first embodiment, but the spacer 14 can be provided on the inner barrel 4 side in the second to eighth embodiments.

  Even when the spacer 14 is provided on the inner barrel 4 side as in the present embodiment, the fan-shaped segment can be urged and restrained outward in the turbine radial direction by the thermal expansion of the spacer 14. Similar effects can be obtained.

  Needless to say, the first to ninth embodiments described above can be arbitrarily combined.

DESCRIPTION OF SYMBOLS 1 Turbine casing 2b Compressed air 3 Compressor exit blade part 4 Inner barrel 5, 5A, B Outer ring 5Aa, Ba Main body part 5Ab, Bb Fitting part 6 Stator blade 7 Inner ring 8 Outer ring hook 11 Concave groove 12 Gap 13 Hook Grooves 14 and 14A Spacer 17 Body portion 18 Fitting portion 19 Slit 20 Surface hardened layer 21 Slit 22 Elastic member 55 Rotating shaft 57 Air compressor AC Wear portion

Claims (10)

An outlet cascade of a gas turbine compressor,
An outer ring that is inserted and fitted into the inner circumferential portion of the turbine casing in the circumferential direction of the turbine;
An inner ring that is inserted and fitted into the recess of the inner barrel surrounding the turbine rotation shaft in the turbine circumferential direction;
A plurality of stationary blades having a tip side connected to the outer ring and a root side connected to the inner ring;
An outlet blade row of a gas turbine compressor, comprising a spacer provided between the turbine casing and the outer ring and formed of a material having a higher coefficient of thermal expansion than the outer ring. An outlet cascade of a gas turbine compressor,
An outer ring that is inserted and fitted into the inner circumferential portion of the turbine casing in the circumferential direction of the turbine;
An inner ring that is inserted and fitted into the recess of the inner barrel surrounding the turbine rotation shaft in the turbine circumferential direction;
A plurality of stationary blades having a tip side connected to the outer ring and a root side connected to the inner ring;
An outlet cascade of a gas turbine compressor, comprising: a spacer provided between the inner barrel and the inner ring and formed of a material having a higher thermal expansion coefficient than the inner ring.   The outlet cascade of the gas turbine compressor according to claim 1 or 2, wherein the spacer is formed in a ring shape.   The outlet cascade of the gas turbine compressor according to claim 1 or 2, wherein the spacer is partially disposed in a circumferential direction of the turbine.   The outlet blade row of the gas turbine compressor according to claim 1 or 2, wherein the outer ring is divided into a main body portion connected to the stationary blade and a fitting portion fitted to the turbine casing. The outlet cascade of the gas turbine compressor.   The outlet blade row of the gas turbine compressor according to claim 1 or 2, wherein the inner ring is divided into a main body portion connected to the stationary blade and a fitting portion fitted to the inner barrel. The outlet cascade of the gas turbine compressor.   3. The gas turbine compressor outlet cascade according to claim 1, wherein the inner ring has a slit extending in the circumferential direction and having a depth in the turbine radial direction. 4. Outlet cascade.   3. The gas turbine according to claim 1, wherein the outer ring and the inner ring are subjected to a surface hardening process on each facing portion of the turbine casing and the inner barrel. 4. Compressor outlet cascade.   3. The gas turbine compressor according to claim 1, wherein the outer ring has a slit extending in the turbine radial direction and having a depth in the turbine radial direction. Outlet cascade.   The outlet cascade of the gas turbine compressor according to claim 1 or 2, wherein an elastic member is disposed between parts adjacent in the circumferential direction of the inner ring.

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