JP2009013837A - Gas turbine facility - Google Patents

Abstract

<P>PROBLEM TO BE SOLVED: To provide a gas turbine facility with which designing of the gas turbine facility and disassembling/assembling work can be easily performed. <P>SOLUTION: In the gas turbine facility, a high pressure gas turbine and a low pressure gas turbine are axially arranged inside turbine casings 5, 5A, 5B, and a working gas flow passage 13 for leading working gas from the high pressure side gas turbine to the low pressure gas turbine is formed between the high pressure gas turbine and the low pressure gas turbine. The working gas flow passage 13 is supported by a component of an initial stage stationary blade 6 of the low pressure gas turbine. <P>COPYRIGHT: (C)2009,JPO&INPIT

Description

  The present invention relates to a gas turbine facility in which a high-pressure gas turbine and a low-pressure gas turbine are arranged in an axial direction, and in particular, operates from a high-pressure gas turbine to a low-pressure gas turbine between an axially adjacent high-pressure gas turbine and low-pressure gas turbine. The present invention relates to a gas turbine facility having a working gas flow path for guiding gas.

  In general, in a gas turbine facility in which a high-pressure gas turbine and a low-pressure gas turbine are arranged in the axial direction, as shown in Patent Document 1, bearings that support a rotating portion are installed in a high-pressure gas turbine and a low-pressure gas turbine, respectively. Therefore, there is a relatively large axial space in the adjacent portion. In order to guide the working gas from the high-pressure gas turbine to the low-pressure gas turbine through such a space, as shown in Patent Document 2, the working gas guides the working gas between the high-pressure gas turbine and the low-pressure gas turbine. A flow path (intermediate duct) is provided.

JP 2004-263623 A (FIG. 1) Japanese Patent Laying-Open No. 2005-9440 (FIG. 5)

  Since the gas turbine equipment shown in Patent Document 2 has a working gas flow path (intermediate duct) provided independently between the high-pressure gas turbine and the low-pressure gas turbine, it is necessary to consider disassembly during assembly and maintenance inspection. Thus, the turbine casing facing the working gas flow path is divided in the axial direction. As a result, the number of parts increases due to the presence of the turbine casing and the independent working gas flow path supported by the turbine casing, which complicates the design and disassembly and assembly work of the gas turbine equipment.

  An object of the present invention is to provide a gas turbine facility that can easily design and disassemble / assemble the gas turbine facility.

  In order to achieve the above object, the present invention arranges a high-pressure gas turbine and a low-pressure gas turbine in a turbine casing in an axial direction, and a low-pressure gas turbine from the high-pressure gas turbine side between the high-pressure gas turbine and the low-pressure gas turbine. In the gas turbine equipment in which the working gas flow path for guiding the working gas to the side is formed, the working gas flow path is supported by the constituent member of the first stage stationary blade of the low pressure gas turbine.

  In this way, by supporting the working gas flow path on the component of the first stage stationary blade of the low-pressure gas turbine, a dedicated turbine casing that supports the working gas flow path becomes unnecessary, and as a result, the number of parts can be reduced. It is possible to easily design and disassemble / assemble the gas turbine equipment.

  A gas turbine facility according to a first embodiment of the present invention will be described below with reference to FIGS.

  In the gas turbine equipment, a rotor 1 of a high-pressure gas turbine and a rotor 2 of a low-pressure gas turbine are arranged adjacent to each other in the axial direction, and are independently rotatably supported by bearings (not shown). The rotor 1 is attached with a final stage moving blade 3 positioned on the most downstream side of the high pressure gas turbine, and the rotor 2 is mounted with a first stage moving blade 4 positioned on the most upstream side of the low pressure gas turbine. The high-pressure gas turbine and the low-pressure gas turbine including the final stage moving blade 3 and the first stage moving blade 4 are covered with a turbine casing 5 that can be divided horizontally in the vertical direction.

  In order to introduce the high-temperature combustion gas G flowing out from the final stage moving blade 3 of the high pressure gas turbine into the first stage moving blade 4 of the low pressure gas turbine at an optimum angle, the first stage stationary blade 6 of the low pressure gas turbine is Arranged just before. A plurality of first stage stationary blades 6 are installed at predetermined intervals in the circumferential direction, and an outer ring 7 is provided on the outer diameter side and an inner ring 8 is provided on the inner diameter side. The outer ring 7 and the inner ring 8 extend to the high-pressure gas turbine side with respect to the first stage stationary blade 6 and are formed to be concentric with the rotors 1 and 2. The outer ring 7 has an engagement mechanism such as an engagement groove and a hook provided on casing shrouds 9 and 10 provided at both ends in the axial direction on the inner peripheral side of the turbine casing 5 facing the final stage moving blade 3 and the first stage moving blade 4. Is held through. Therefore, as a result, the outer ring 7 is held in the turbine casing 5, and the inner ring 8 is also supported by the turbine casing 5 via the first stage stationary blade 6 by holding the outer ring 7 in the turbine casing 5. It will be.

  Further, a diaphragm 11 is attached to the inner diameter side of the inner ring 8 by an engagement mechanism such as an engagement groove and a hook, and a partition wall 12 is provided on the inner diameter side of the diaphragm 11 so that a high pressure gas turbine and a low pressure gas turbine are provided. Partitioning. Accordingly, the combustion gas G from the high pressure gas turbine exits the final stage moving blade 3 and then passes through the working gas flow path 13 constituted by the outer ring 7 and the inner ring 8 of the first stage stationary blade 6, so The blade 6 is reached and introduced into the first stage blade 4.

  In the above description, the working gas flow path 13 is formed by the outer ring 7 and the inner ring 8, but the working gas flow path 13 is composed of a concentric outer peripheral wall and an inner peripheral wall. The inner peripheral wall may be extended to the low-pressure gas turbine side and connected to the outer diameter portion and the inner diameter portion of the first stage stationary blade 6, and the outer ring 7 and the inner ring 8 of the first stage stationary blade 6 may be combined. In other words, the working gas flow path 13 is constituted by concentric outer peripheral walls and inner peripheral walls, which are connected to the first stage stationary blade 6 itself or the outer ring 7 and the inner ring 8 which are constituent members of the first stage stationary blade 6. Alternatively, the outer ring 7 and the inner ring 8 of the first stage stationary blade 6 may be extended to form the outer peripheral wall and inner peripheral wall of the working gas flow path 13.

  By the way, the outer ring 7 and the inner ring 8 have a segment structure divided into a plurality of parts in the circumferential direction including the first stage stationary blade 6. When the inner ring 8 is viewed as an example, as shown in FIG. 2, seal grooves 8G are formed on the opposing surfaces of the adjacent inner ring segments 8A and 8B, and a plate-like seal key 8S is fitted into these seal grooves 8G. I wear it to close the gap.

  Further, a minute gap 14A is provided on the upstream side of the inner ring 8 so as not to be in contact with the final stage moving blade 3 and the downstream side is extended to a position just before the first stage moving blade 4 to be in contact. A small gap 14B is provided so as not to occur.

  In the above configuration, the combustion gas G flowing out from the final stage moving blade 3 of the high pressure gas turbine is guided to the working gas flow path 13 constituted by the outer ring 7 and the inner ring 8 which are constituent members of the first stage stationary blade 6, and means. After being rectified by the stationary blade 6, it is introduced into the first stage blade 4 of the low-pressure gas turbine.

  When the combustion gas G passes through the working gas flow path 13, the outer ring 7 suppresses the leakage of the combustion gas G toward the turbine casing 5, so that the high-temperature combustion gas G does not reach the turbine casing 5. . As a result, the temperature drop of the combustion gas G can be prevented, so that the heat efficiency can be prevented from being lowered, and the turbine casing 5 can be prevented from being overheated.

  In addition, since the working gas flow path 13 has an integral structure with the first stage stationary blade 6 and has a segment structure, the segment dividing surface is made to coincide with the dividing surface of the turbine casing 5, so that the turbine casing can be used for disassembly. 5 can be decomposed simultaneously. Then, after the turbine casing 5 is disassembled, the disassembly work can be easily performed by disassembling the outer ring 7 and the inner ring 8 including the first stage stationary blade 6 having the segment structure. Therefore, the assembling work may be performed in the reverse order of the disassembling work and can be easily performed.

  In addition, since the working gas flow path 13 is configured by the outer ring 7 and the inner ring 8 which are constituent members of the first stage stationary blade 6 as described above, the present embodiment is dedicated to support the working gas flow path 13. This eliminates the need for additional installation of the turbine casing. As a result, the number of parts can be reduced, and the design and disassembly / assembly of the gas turbine equipment can be easily performed.

  Further, since the working gas flow path 13 is constituted by the outer ring 7 and the inner ring 8 which are constituent members of the first stage stationary blade, the entire flow path through which the working gas G flows compared to the conventional case where the working gas flow path 13 is separately provided. Thus, the turbulent flow of the working gas G can be reduced, and high aerodynamic performance can be obtained.

  Furthermore, since the outer ring 7 and the inner ring 8 constituting the working gas flow path 13 have a segment structure, it is assumed that the working gas flow path 13 flows immediately after the last stage blade 3 and immediately before the first stage blade 4. Even in the case where the path diameter changes greatly, in this embodiment, only the curvature of the segmented curved peripheral wall piece and the width in the circumferential direction are changed as compared with the outer ring 7 and the inner ring 8 integrally formed in an annular shape. It can be easily dealt with and the machining operation can be easily performed. In addition, since the outer ring 7 and the inner ring 8 are formed by gathering the peripheral wall pieces, the thickness of the peripheral wall piece can be increased compared to the outer ring 7 and the inner ring 8 that are integrally formed in an annular shape. And the thickness of the inner ring 8 can be increased. Therefore, the life of the outer ring 7 and the inner ring 8 due to high temperature oxidation, high temperature corrosion, and thinning due to erosion can be extended as compared with the outer ring 7 and the inner ring 8 integrally formed in an annular shape.

  Furthermore, the outer ring 7 and the inner ring 8 are made of the same material as the first stage stationary blade 6 that has the same high load and high stress as the high temperature strength, thereby obtaining a highly reliable working gas channel 13 having a long life. Can do.

  Next, a second embodiment of the gas turbine equipment according to the present invention will be described with reference to FIG. The same reference numerals as those in FIGS. 1 and 2 indicate the same components, and detailed description thereof will not be repeated.

  The configuration different from the first embodiment is that the turbine casing is divided into a high pressure side turbine casing 5A and a low pressure side turbine casing 5B. The division position is a position facing immediately after the final stage rotor blade 3, and flanges 5FA and 5FB are provided on the high pressure side turbine casing 5A and the low pressure side turbine casing 5B at that position and fastened by fastening means such as bolts and nuts. It is.

  A casing shroud 15 is provided on the inner diameter side where the flange 5FB of the low-pressure side turbine casing 5B is formed by the same means as in the first embodiment, and the outer ring 7 is attached to the casing shroud 15 and the casing shroud 10 in the first position. It was attached in the same manner as in the first embodiment.

  By configuring in this way, the outer ring 7 has a shorter axial length than the inner ring 8, but the inner peripheral surface 15S of the casing shroud 15 and the inner peripheral surface of the outer ring 7 are continuous without any step, so that the first A working gas flow path 13 having substantially the same size as that of the first embodiment and having no step is formed.

  By configuring in this way, the same effects as those of the first embodiment can be obtained, and only the high-pressure turbine casing 5A or only the low-pressure turbine casing 5B can be disassembled independently, improving the maintenance and inspection performance. Can be made.

  However, when the end portion 8E of the inner ring 8 facing the final stage moving blade 3 is positioned slightly downstream from the dividing surface of the high-pressure turbine casing 5A and the low-pressure turbine casing 5B, respectively, In addition, it is necessary to avoid the end 8E of the inner ring 8 from coming into contact with the components of the high-pressure gas turbine.

  FIG. 4 shows a third embodiment in which a cooling structure is added to the first embodiment of FIG. 1, and the same reference numerals as those in FIGS. Is omitted.

  In the present embodiment, the sealed space surrounded by the turbine casing 5, the casing shrouds 9, 10 and the outer ring 7 is used as the first cooling medium chamber 16, and the impingement cover 17 provided with a plurality of through holes 17 </ b> H is used as the outer ring 7. Cooling medium supply holes 5H, 9H, and 10H are provided in the turbine casing 5 and the casing shrouds 9 and 10, respectively, that are installed on the outer peripheral side through a space and supply the cooling medium to the first cooling medium chamber 16. The cooling medium supply holes 5H, 9H, and 10H are not necessarily provided depending on the specifications and models of the gas turbine equipment, and may be provided selectively.

  The outer ring 7 has a plurality of through-holes 7H1 penetrating into the working gas passage 13 and a through-hole 7H2 penetrating through the first stage stationary blade 6 in the radial direction and communicating with the hollow portion 6H constituting the cooling medium passage. Forming.

  On the other hand, the sealed space surrounded by the inner ring 8 and the diaphragm 11 is used as the second cooling medium chamber 18, and the impingement cover 19 provided with a plurality of through holes 19H is installed on the inner peripheral side of the inner ring 8 through the space. In order to supply the cooling medium to the second cooling medium chamber 18 via the hollow portion 6H of the first stage stationary blade 6, the inner ring 8 is provided with a cooling medium supply hole 8H2.

  Further, the inner ring 8 is formed with a plurality of through holes 8H1 penetrating into the working gas flow path 13, and the diaphragm 11 has a through hole 11H1 for discharging the cooling medium to the final stage blade 3 and the first stage blade 4. A through hole 11H2 for discharging a cooling medium is formed on the side.

  In the above configuration, each part can be cooled in addition to the same effects as those of the first embodiment. Specifically, when, for example, cooling air, which is a cooling medium, is supplied to the first cooling medium chamber 16 from the outside via all or part of the cooling medium supply holes 5H, 9H, 10H, the impingement cover 17 A plurality of through-holes 17H are introduced into the outer peripheral portion of the outer ring 7 and a part thereof is discharged from the through-hole 7H1 into the working gas flow path 13 to cool the outer ring 7 by convection cooling and film cooling. The remaining part is introduced into the second cooling medium chamber 18 on the inner diameter side of the inner ring 8 through the through hole 7H2 and the hollow portion 6H of the first stage stationary blade 6. A part of the cooling air introduced into the second cooling medium chamber 18 is discharged into the working gas flow path 13 from the plurality of through holes 19H of the impingement cover 19, and cools the inner ring 8 by convection cooling and film cooling. The remaining part of the air is discharged from the through holes 11H1 and 11H2 to cool the vicinity of the final stage moving blade 3 and the first stage moving blade 4.

  Thus, according to the present embodiment, each part can be cooled in addition to the same effects as those of the first embodiment.

  FIG. 5 is based on the third embodiment of FIG. 4 and is mainly composed of the cooling structure of the rotors 1 and 2. The same reference numerals as those in FIGS. 1 and 4 indicate the same components. The detailed description of is omitted.

  The cooling medium supplied to the first cooling medium chamber 16 formed by the turbine casing 5, the casing shrouds 9 and 10, and the outer ring 7 is passed through the inner ring 8 and the diaphragm 11 via the hollow portion 6 </ b> H of the first stage stationary blade 6. The diaphragm 11 is connected to the two partition walls 12A and 12B. The diaphragm 11 is connected to the two partition walls 12A and 12B. The two partition walls 12A and 12B are formed with cooling medium supply paths 20A and 20B, respectively. Through-holes 11HA and 11HB for flowing a cooling medium to 20B are formed in the diaphragm 11.

  With this configuration, the cooling medium in the first cooling medium chamber 16 passes through the hollow portion 6H of the first stage stationary blade 6 serving as a cooling medium flow path, and the second cooling medium chamber 18 on the inner ring 8 side. Led to. The cooling medium guided to the second cooling medium chamber 18 is press-fitted into the cooling medium flow paths 20A and 20B, and is pumped to the inner diameter side. Since the coolant jets 20AH and 20BH are opened in the axial direction on the inner diameter side of the coolant flow paths 20A and 20B so as to face the rotors 1 and 2 facing each other, they are blown out from the coolant jets 20AH and 20BH. A part of the cooling medium proceeds as it is and enters the center holes 1R and 2R formed in the inner diameter portions of the rotors 1 and 2 to cool the rotors 1 and 2 from the inner diameter side, and the remaining part proceeds in the outer diameter direction to the rotor. The surface is cooled.

  Therefore, according to the present embodiment, the cooling effect of the rotors 1 and 2 can be enhanced in addition to the effects equivalent to those of the first embodiment.

  By the way, although each embodiment may be implemented independently, it is needless to say that each configuration of each embodiment may be appropriately incorporated and combined according to the specifications and applications of the gas turbine equipment. is there.

  Further, in the third embodiment, the impingement covers 17 and 19 may be provided in the axial direction over almost the entire length of the working gas flow path 13 or partially provided near the high-pressure gas turbine. Also good. Furthermore, the impingement covers 17 and 19 themselves are not necessarily provided if it is desired to reduce the number of parts.

  Further, it is important that the working gas flow path 13 is configured by the outer ring 7 and the inner ring 8 of the first stage stationary blade 6 in each of the above embodiments, but the working gas is a member different from the constituent members of the first stage stationary blade 6. Of course, the outer peripheral wall and the inner peripheral wall of the flow path may be formed, and the inner and outer peripheral walls may be connected to the outer ring 7 and the inner ring 8 which are constituent members of the first stage stationary blade 6.

The principal part vertical side view which shows 1st Embodiment of the gas turbine equipment by this invention. The expanded sectional view which follows the AA line of FIG. The principal part vertical side view which shows 2nd Embodiment of the gas turbine equipment by this invention. The principal part vertical side view which shows 3rd Embodiment of the gas turbine equipment by this invention. The principal part vertical side view which shows 4th Embodiment of the gas turbine equipment by this invention.

Explanation of symbols

  DESCRIPTION OF SYMBOLS 1, 2 ... Rotor, 1R, 2R ... Center hole, 3 ... Final stage moving blade, 4 ... First stage moving blade, 5 ... Turbine casing, 6 ... First stage stationary blade, 7 ... Outer ring, 8 ... Inner ring, 8A, 8B ... Inner ring Segment, 8S ... Seal key, 9, 10, 15 ... Casing shroud, 11 ... Diaphragm, 12, 12A, 12B ... Partition, 13 ... Working gas flow path, 14A, 14B ... Micro gap, 16, 18 ... Sealed space, 17 , 19 ... Impingement cover, 7H1, 7H2, 11H1, 11H2, 17H, 19H ... Through-hole, 20A, 20B ... Cooling medium flow path.

Claims (12)

  A high-pressure gas turbine and a low-pressure gas turbine are arranged in the turbine casing in the axial direction, and a working gas passage for guiding the working gas from the high-pressure gas turbine side to the low-pressure gas turbine side is provided between the high-pressure gas turbine and the low-pressure gas turbine. In the formed gas turbine equipment, the working gas flow path is supported by a constituent member of a first stage stationary blade of a low-pressure gas turbine.   A high-pressure gas turbine and a low-pressure gas turbine are arranged in the turbine casing in the axial direction, and a working gas passage for guiding the working gas from the high-pressure gas turbine side to the low-pressure gas turbine side is provided between the high-pressure gas turbine and the low-pressure gas turbine. In the formed gas turbine equipment, the working gas flow path has concentric inner and outer peripheral walls, and these inner and outer peripheral walls are connected to an inner ring and an outer diameter provided on the inner diameter side of the first stage stationary blade of the low-pressure gas turbine. Gas turbine equipment, characterized by being supported by an outer ring provided on the side.   A high pressure gas turbine and a low pressure gas turbine are arranged in the casing in the axial direction, and a working gas flow path is formed between the high pressure gas turbine and the low pressure gas turbine to guide the working gas from the high pressure gas turbine side to the low pressure gas turbine side. In the gas turbine equipment, the first stage stationary blade of the low-pressure gas turbine has an outer ring provided on the outer diameter side and an inner ring provided on the inner diameter side, and the inner gas and the outer ring constitute the working gas flow path. A gas turbine facility characterized in that concentric inner and outer peripheral walls are formed.   Formed by a high-pressure gas turbine and a low-pressure gas turbine arranged along the axial direction, a turbine casing covering these gas turbines, and an inner peripheral wall and an outer peripheral wall arranged concentrically between the high-pressure gas turbine and the low-pressure gas turbine A working gas flow path for guiding the working gas from the high pressure gas turbine side to the low pressure gas turbine side, and a constituent member of the first stage stationary vane of the low pressure gas turbine supporting the inner peripheral wall and the outer peripheral wall of the working gas flow path. Gas turbine equipment.   High pressure gas turbines and low pressure gas turbines arranged along the axial direction, turbine casings covering these gas turbines, and working gas from the high pressure gas turbine side to the low pressure gas turbine side between the high pressure gas turbine and the low pressure gas turbine. A first working gas passage formed between the outer peripheral wall and the turbine casing; a working gas passage for guiding; a constituent member of the first stage stationary blade of the low-pressure gas turbine supporting the inner and outer peripheral walls constituting the working gas passage; Cooling medium chamber, a second cooling medium chamber formed on the inner diameter side of the inner peripheral wall, and cooling that guides the cooling medium from the first cooling medium quality to the second cooling medium chamber via the first stage stationary blade. Gas turbine equipment comprising a medium flow path.   The first stage stationary blade is composed of a plurality of stationary blade segments divided in the circumferential direction, and the inner circumferential wall and the outer circumferential wall constituting the working gas flow path are also composed of the same number of inner circumferential wall segments and outer circumferential wall segments as the stationary blade segments. The gas turbine equipment according to claim 5, wherein the gas turbine equipment is configured.   The gas turbine equipment according to claim 6, wherein the inner peripheral wall segment and the outer peripheral wall segment close a gap between adjacent segments with a seal key.   The outer peripheral wall is provided with a through hole for discharging the cooling medium into the working gas flow path, and the inner peripheral wall is provided with a through hole for discharging the cooling medium into the working gas flow path. The gas turbine equipment according to claim 5.   The gas turbine equipment according to claim 5, wherein the second cooling medium chamber is provided with a through hole that discharges the cooling medium on an inner diameter side of the inner peripheral wall.   The gas turbine according to claim 5, wherein a cooling medium supply path is formed from the second cooling medium chamber to guide the cooling medium to a rotor center portion side of the high pressure gas turbine and the low pressure gas turbine. Facility.   The axially opposite ends of the outer peripheral wall are supported by the turbine casing via a casing shroud, and the inner peripheral wall is supported by a partition wall that partitions the high pressure gas turbine and the low pressure gas turbine. Item 10. The gas turbine equipment according to any one of Items 2 to 9.   10. The turbine casing according to claim 1, wherein the turbine casing is divided into a high-pressure turbine casing and a low-pressure turbine casing at a position downstream of the downstream blades of the high-pressure turbine. The gas turbine equipment described.

Images