Turbine seal system and method of assembly thereof

Abstract

 A system includes a multi-stage turbine (22). The multi-stage turbine (22) includes a first turbine stage (62) including a first wheel (66) having a plurality of first blade segments (68) spaced circumferentially about the first wheel (66), a second turbine stage (64) including a second wheel (74) having a plurality of second blade segments (76) spaced circumferentially about the second wheel (74), and an interstage seal (40) extending axially between the first (62) and second (64) turbine stages. The interstage seal (40) is configured to be installed or removed while the first (66) and second (74) wheels remain in place in the respective first (62) and second (64) turbine stages.

Description

BACKGROUND OF THE INVENTION

[0001] The subject matter disclosed herein relates to gas turbines, and more specifically, to seals within turbines.

[0002] In general, gas turbine engines combust a mixture of compressed air and fuel to produce hot combustion gases. The combustion gases may flow through one or more turbine stages to generate power for a load and/or compressor. A pressure drop may occur between stages, which may allow leakage flow of a fluid, such as combustion gases, through unintended paths. Seals may be disposed between the stages to reduce fluid leakage between stages. Unfortunately, the seals may be subject to stresses, such as thermal stresses, which may bias the seals in axial and/or radial directions thereby reducing effectiveness of the seals. For example, seal deflection may increase the possibility of a rub condition between stationary and rotating components.

BRIEF DESCRIPTION OF THE INVENTION

[0003] Certain embodiments commensurate in scope with the originally claimed invention are summarized below. These embodiments are not intended to limit the scope of the claimed invention, but rather these embodiments are intended only to provide a brief summary of possible forms of the invention. Indeed, the invention may encompass a variety of forms that may be similar to or different from the embodiments set forth below.

[0004] In a first aspect, the present invention resides in a system including a multi-stage turbine. The multi-stage turbine includes a first turbine stage including a first wheel having a plurality of first blade segments spaced circumferentially about the first wheel, a second turbine stage including a second wheel having a plurality of second blade segments spaced circumferentially about the second wheel, and an interstage seal extending axially between the first and second turbine stages. The interstage seal is configured to be installed or removed while the first and second wheels remain in place in the respective first and second turbine stages.

[0005] In a second aspect, the invention resides in a method including positioning a first recessed portion of an interstage seal about a first wheel rim of a turbomachine, pivoting a second recessed portion of the interstage seal toward an axial axis of the turbomachine, and moving the interstage seal along the axial axis toward a second wheel rim of the turbomachine to position the second recessed portion about the second wheel rim.

BRIEF DESCRIPTION OF THE DRAWINGS

[0006] Embodiments of the present invention will now be described, by way of example only, with reference to the accompanying drawings in which:

FIG. 1 is a schematic flow diagram of an embodiment of a gas turbine engine that may employ turbine seals;

FIG. 2 is a cross-sectional side view of an embodiment of the gas turbine engine of FIG. 1 taken along the longitudinal axis;

FIG. 3 is a partial cross-sectional side view of the gas turbine engine of FIG. 2 illustrating an embodiment of a seal structure between turbine stages;

FIG. 4 is a partial cross-sectional side view of the gas turbine engine of FIG. 2 illustrating an embodiment of a seal structure being pivoted between adjacent stages;

FIG. 5 is a partial cross-sectional side view of the gas turbine engine of FIG. 2 illustrating an embodiment of a seal structure being moved along the longitudinal axis between adjacent stages;

FIG. 6 is a partial cross-sectional side view of the gas turbine engine of FIG. 2 illustrating an embodiment of a seal structure being installed between adjacent stages;

FIG. 7 is a front perspective view of an embodiment of a seal structure;

FIG. 8 is a rear perspective view of an embodiment of a seal structure;

FIG. 9 is a front view of an embodiment of a seal structure; and

FIG. 10 is a side view of an embodiment of a seal structure.

DETAILED DESCRIPTION OF THE INVENTION

[0007] One or more specific embodiments of the present invention will be described below. In an effort to provide a concise description of these embodiments, all features of an actual implementation may not be described in the specification. It should be appreciated that in the development of any such actual implementation, as in any engineering or design project, numerous implementation-specific decisions must be made to achieve the developers' specific goals, such as compliance with system-related and business-related constraints, which may vary from one implementation to another. Moreover, it should be appreciated that such a development effort might be complex and time consuming, but would nevertheless be a routine undertaking of design, fabrication, and manufacture for those of ordinary skill having the benefit of this disclosure.

[0008] When introducing elements of various embodiments of the present invention, the articles "a," "an," "the," and "said" are intended to mean that there are one or more of the elements. The terms "comprising," "including," and "having" are intended to be inclusive and mean that there may be additional elements other than the listed elements.

[0009] The present disclosure is directed to gas turbine engines that include interstage seals, wherein each interstage seal includes features to seal an interstage gap without the use of additional components, such as spacer wheels. Thus, gas turbine engines that include such interstage seals may be less costly than engines using spacer wheels. For example, the gas turbine engine may include a first turbine stage that includes a first wheel that has a plurality of first blade segments spaced circumferentially about the first wheel, and a second turbine stage that includes a second wheel having a plurality of second blade segments spaced circumferentially about the second wheel. The interstage seal may extend axially between the first and second turbine stages to seal the interstage gap. In addition, embodiments of the interstage seal may be installed and removed without disassembling a rotor of the gas turbine engine. For example, the interstage seal may be configured to be installed or removed while the first and second wheels remain in place in the respective first and second turbine stages. Thus, if only the interstage seal is replaced, the rotor of the gas turbine engine need not be disturbed, thereby potentially reducing maintenance time, complexity, and/or cost. In further embodiments, the interstage seal may include an inclined support rib that is configured to enable the interstage seal to pivot toward and away from an axial axis of the gas turbine engine without removal of the first wheel or the second wheel. In other words, pivoting of the interstage seal may enable the interstage seal to be replaced without disturbing the rotor assembly. In other embodiments, a recessed portion of the interstage seal may be configured to enable the pivoting of the interstage seal.

[0010] FIG. 1 is a block diagram of an exemplary system 10 including a gas turbine engine 12 that may employ interstage seals configured to be installed or removed without rotor disassembly, as described in detail below. In certain embodiments, the system 10 may include an aircraft, a watercraft, a locomotive, a power generation system, or combinations thereof. The illustrated gas turbine engine 12 includes an air intake section 16, a compressor 18, a combustor section 20, a turbine 22, and an exhaust section 24. The turbine 22 is coupled to the compressor 18 via a shaft 26.

[0011] As indicated by the arrows, air may enter the gas turbine engine 12 through the intake section 16 and flow into the compressor 18, which compresses the air prior to entry into the combustor section 20. The illustrated combustor section 20 includes a combustor housing 28 disposed concentrically or annularly about the shaft 26 between the compressor 18 and the turbine 22. The compressed air from the compressor 18 enters combustors 30, where the compressed air may mix and combust with fuel within the combustors 30 to drive the turbine 22.

[0012] From the combustor section 20, the hot combustion gases flow through the turbine 22, driving the compressor 18 via the shaft 26. For example, the combustion gases may apply motive forces to turbine rotor blades within the turbine 22 to rotate the shaft 26. After flowing through the turbine 22, the hot combustion gases may exit the gas turbine engine 12 through the exhaust section 24. As discussed below, the turbine 22 may include a plurality of interstage seals, which may be installed or removed while rotating components of the turbine 22, such as wheels, remain in place. Thus, maintenance affecting the interstage seals may be performed without complete disassembly of the turbine 22.

[0013] FIG. 2 is a cross-sectional side view of an embodiment of the gas turbine engine 12 of FIG. 1 taken along the longitudinal axis 32. As depicted, the gas turbine 22 includes three separate stages 34. Each stage 34 includes a set of blades 36 coupled to a rotor wheel 38 that may be rotatably attached to the shaft 26 (FIG. 1). The blades 36 extend radially outward from the rotor wheels 38 and are partially disposed within the path of the hot combustion gases. Seals 40 extend between and are supported by adjacent rotor wheels 38. As discussed below, the seals 40 may include recessed portions that fit about adjacent wheels 38 for support. The recessed portions may be configured to enable the seals 40 to pivot toward and away from the longitudinal axis 32 during installation or removal. Thus, the seals 40 may be installed or removed while the rotor wheels 38 remain in place in the gas turbine engine 12. In addition, the seals 40 may provide for improved cooling of the stages 34. Although the gas turbine 22 is illustrated as a three-stage turbine, the seals 40 described herein may be employed in any suitable type of turbine with any number of stages and shafts. For example, the seals 40 may be included in a single stage gas turbine, in a dual turbine system that includes a low-pressure turbine and a high-pressure turbine, or in a steam turbine. Further, the seals 40 described herein may also be employed in a rotary compressor, such as the compressor 18 illustrated in FIG. 1. The seals 40 may be made from various high-temperature alloys, such as, but not limited to, nickel based alloys.

[0014] As described above with respect to FIG. 1, air enters through the air intake section 16 and is compressed by the compressor 18. The compressed air from the compressor 18 is then directed into the combustor section 20 where the compressed air is mixed with fuel. The mixture of compressed air and fuel is generally burned within the combustor section 20 to generate high-temperature, high-pressure combustion gases, which are used to generate torque within the turbine 22. Specifically, the combustion gases apply motive forces to the blades 36 to turn the wheels 38. In certain embodiments, a pressure drop may occur at each stage 34 of the turbine 22, which may allow gas leakage flow through unintended paths. For example, the hot combustion gases may leak into the interstage volume between turbine wheels 38, which may place thermal stresses on the turbine components. In certain embodiments, the interstage volume may be cooled by discharge air bled from the compressor or provided by another source. However, flow of hot combustion gases into the interstage volume may abate the cooling effects. Accordingly, the seals 40 may be disposed between adjacent wheels 38 to seal and enclose the interstage volume from the hot combustion gases. In addition, the seals 40 may be configured to direct a cooling fluid to the interstage volume or from the interstage volume toward the blades 36.

[0015] FIG. 3 is a cross-sectional side view of an embodiment of a pair of adjacent rotor stages 34 shown in FIG. 2. In the following discussion, reference may be made to an axial direction or axis 50, a radial direction or axis 52, and a circumferential direction or axis 54, relative to the longitudinal axis 32 of the gas turbine engine 12. Hot fluids, such as hot combustion gases or steam, with a flow path 56 (illustrated generally by an arrow) enters at an upstream side 58 and exits at a downstream side 60. For illustrative purposes, only a portion of the stages 34 are illustrated in FIG. 3. Specifically, a first turbine stage 62 is shown near the upstream side 58 and a second turbine stage 64 is shown near the downstream side 60. The first turbine stage 62 includes a first wheel 66 with a plurality of first blade segments 68 extending radially outward 52 from a first wheel post portion 70 of the first wheel 66. The first wheel post portion 70 is disposed along the circumference of the first wheel 66 and includes slots 72 (e.g., axial dovetail slots) for retaining lower segments (e.g., axial dovetail tabs 73) of the first blade segments 68. Similarly, the second turbine stage 64 includes a second wheel 74 with a plurality of second blade segments 76 extending radially outward 52 from a second wheel post portion 78 of the second wheel 74. The second wheel post portion 78 is disposed along the circumference of the second wheel 74 and includes slots 80 (e.g., axial dovetail slots) for retaining lower segments (e.g., axial dovetail tabs 81) of the plurality of second blade segments 76. In certain embodiments, approximately 50 to 150 first and second blade segments 68 and 76 may be mounted and spaced circumferentially 54 around the first and second wheels 66 and 74 and a corresponding axis of rotation (extending generally in the direction indicated by arrow 50). In further embodiments, methods other than the slots and tabs described above may be used to couple the first and second blade segments 68 and 76 to the first and second wheels 66 and 74.

[0016] The interstage seal 40 extends between the first and second adjacent wheels 66 and 74 and is mechanically supported by the first and second turbine stages 62 and 64. As described in detail below, an annular interstage seal assembly 41 (as shown in FIG. 9) may include a plurality of interstage seal segments 40 disposed about the longitudinal axis 32 of the gas turbine engine 12. In other words, the interstage seal assembly 41 is a segmented annular seal assembly. The interstage seal segment 40 includes an outer bridge portion 82, or axial beam, disposed near the first and second pluralities of blade segments 68 and 76. The outer bridge portion 82 is a structure that provides support for the interstage seal 40. The interstage seal segment 40 also includes an inner bridge portion 84 disposed near the first and second wheels 66 and 74. The inner bridge portion 84 also provides support for the inner stage seal 40. In addition, the inner bridge portion 84 may have a catenary shape, e.g., a curved annular shape, configured to provide additional strength to the interstage seal 40. At intermediate locations between the outer bridge portion 82 and the inner bridge portion 84, the interstage seal 40 includes one or more intermediate supports 86, or radial beams, which provide support for the inner stage seal 40 in the radial direction 52. As illustrated in FIG. 3, the intermediate supports 86 may be generally aligned with the radial direction 52. In other embodiments, the interstage seal 40 may include three, four, five, six, or more intermediate supports 86. The intermediate supports 86 may be disc-shaped structures that generally taper in the radial direction 52. In other words, the intermediate supports 86 may be thicker near the inner bridge portion 84 than near the outer bridge portion 82.

[0017] The interstage seal 40 may also include an inclined support rib 88, or support beam, that may be disposed between the inner and outer bridge portions 82 and 84. As shown in FIG. 3, the inclined support rib 88 may be inclined with respect to the radial direction 52. The interstage seal 40 may also include an optional inclined support portion 90 to provide additional support for the portion of the outer bridge portion 82 not supported by the inclined support rib 88. In certain embodiments, the inclined support portion 90 may be omitted. In further embodiments, the inclined support portion 90 may have a generally triangular cross sectional shape. Because the inclined support rib 88 may be inclined with respect to the radial direction 52, the inclined support rib 88 may form an outer angle 92 with the outer bridge portion 82 and an inner angle 94 with the inner bridge portion 84. As shown in FIG. 3, the outer and inner angles 92 and 94 may be acute angles of less than approximately 90 degrees. For example, in certain embodiments, the outer and inner angles 92 and 94 may be between approximately 10 to 80 degrees, 20 to 70 degrees, 30 to 60 degrees, or 40 to 50 degrees. In one embodiment, the outer and inner angles 92 and 94 may be less than approximately 75 degrees. As discussed in detail below, the inclined support rib 88 enables the interstage seal 40 to be pivoted during installation and removal from the gas turbine engine 10. Thus, the interstage seal 40 may be installed or removed without removal of the first and second wheels 66 and 74. In addition, in certain embodiments, the interstage support portion 90 may be coupled to the inner bridge portion 84 instead of the outer bridge portion 82.

[0018] Seal cavities 96 may be formed in the interstage seal 40 between the intermediate supports 86. The seal cavities 96 may enable a cooling fluid, such as air, to circulate between the first and second turbine stages 62 and 64 as discussed in detail below. Recessed portions 98 may be formed between the outer and inner bridge portions 82 and 84 near the ends of the interstage seal 40 facing toward the first and second turbine stages 62 and 64. Specifically, the intermediate supports 86 and the inclined support rib 88 may not be located at the ends of the outer and inner bridge portions 82 and 84. Thus, the recessed portions 98 are formed in the spaces surrounded by the intermediate supports 86, the inclined support rib 88, and the outer and inner bridge portions 82 and 84. The seal cavities 96 may have a variety of cross sectional shapes depending on the configuration of the intermediate supports 86 and the inclined support rib 88. For example, the seal cavities 96 may have rectangular, square, triangular, circular, oval, or other suitable cross sectional shapes. Similarly, the recessed portions 98 may have a variety of cross sectional shapes, such as, but not limited to, rectangular, square, triangular, circular, oval, or other suitable shapes. In addition, the inclined support portion 90 may occupy part of the recessed portion 98 adjacent to the inclined support rib 88. In other embodiments, the inclined support portion 90 may be omitted. As discussed in detail below, the recessed portions 98 may at least partially fit over portions of the first and second turbine stages 62 and 64. In other words, portions of the first and second wheels 66 and 74 may extend into the recessed portions 98 to enable pivoted motion of the interstage seal 40 and/or enable installation and removal of interstage seal 40 without removal of the first and second wheels 66 and 74.

[0019] In certain embodiments, a labyrinth seal 100 may be disposed adjacent to the interstage seal 40 and between the first and second turbine stages 62 and 64. The labyrinth seal 100 may be configured to help block axial leakage of the hot combustion gases 56. For example, the labyrinth seal 100 may include an abradable coating 102 on the surface facing toward the interstage seal 40. Correspondingly, the interstage seal 40 may include one or more teeth 104 disposed adjacent to the abradable coating 102. During operation of the gas turbine engine 10, the teeth 104 may be in close proximity to the abradable coating 102 to help block axial leakage of the hot combustion gases 56 between the first and second turbine stages 62 and 64. In response to transient conditions, such as rotor transients, the abradable coating 102 may be configured to partially abrade when in contact with the teeth 104 to help prevent damage to the teeth 104. In other words, the abradable coating 102 may be softer than the teeth 104. In further embodiments, seals other than the labyrinth seal 100 may be used together with the interstage seal 40.

[0020] The portions of the outer bridge portion 82 that extends past the intermediate support 86 and the inclined support rib 88 may be referred to as end portions. Specifically, the outer bridge portion 82 may include a first end portion 106 and a second end portion 108. In certain embodiments, the first and second end portions 106 and 108 may include optional centrifugal seals 110 to help block radial leakage of the hot combustion gases 56. For example, the first and second end portions 106 and 108 may include a recessed slot 111 to engage with the centrifugal seal 110. The seal 110 may include a support rod 112, a curved support piece 114, and a seal rod 116. The support rod 112 of the centrifugal seal 110 may fit in the recessed slot 111. The curved support piece 114 may be attached to the support rod 112. Finally, the seal rod 116 may be attached to the end of the curved support piece 114. When the gas turbine engine 10 is operating, centrifugal forces may cause the seal rod 116 to move away from the interstage seal 40 and toward the surfaces of the first and second turbine stages 62 and 64 facing the interstage seal 40. Thus, the seal rod 116 may be in contact with the first and second blade segments 68 and 76 during operation of the gas turbine engine 10 to help block radial leakage of the hot combustion gases 56. To accommodate the movement of the centrifugal seals 110 during operation of the gas turbine engine 10, small gaps exist between the first and second end portions 106 and 108 of the interstage seal 40 and the first and second turbines stages 62 and 64. By moving toward or away from the interstage seal 40, the centrifugal seals 110 may be able to maintain contact with the first and second turbine stages 62 and 64 even during axial transients that may cause the gaps to increase or decrease during operation of the gas turbine engine 10. In other embodiments, the centrifugal seals 110 may be omitted or seals other than the centrifugal seals 110 may be used at the outer bridge portion 82 to provide for radial sealing.

[0021] In certain embodiments, the second end portion 108 may include a first support feature 118 configured to engage with a second support feature 120 disposed on one or more of the second blade segments 76. For example, the first support feature 118 may be a female alignment portion (e.g., a notch) and the second support feature 120 may be a male alignment portion (e.g., a tab). In other embodiments, the first support feature 118 may be the male alignment portion, and the second support feature 120 may be the female alignment portion. Together, the first and second support features 118 and 120 may help to block radial movement of the interstage seal 40 in the direction 52 toward the axial axis 50 of the gas turbine engine 10 during installation or removal of the interstage seal 40. In addition, the first and second support features 118 and 120 may help to block circumferential movement of the interstage seal 40 in the direction 54 during operation of the gas turbine engine 10. Use of the first and second support features 118 and 120 during installation and removal of the interstage seal 40 is described in detail below.

[0022] The inner bridge portion 84 may also include end portions, specifically, a first end portion 124, and a second end portion 126. The first end portion 124 may be configured to engage with a first wheel rim 128 of the first wheel 66 during operation of the gas turbine engine 10. Specifically, during operation of the gas turbine engine 10, centrifugal forces may move the interstage seal 40 in the radial direction 52 toward the first rim 128. Contact between the first end portion 124 and the first rim 128 may provide an additional seal against radial leakage of the hot combustion gases 56. The first end portion 124 may include an axial stop 130 disposed in the recessed portion 98. The axial stop 130 may be a structure configured to restrict movement of the interstage seal 40 in the axial direction 50 toward the first turbine stage 62. Similarly, the second end portion 126 may be configured to engage with a second wheel rim 132 of the second wheel 74 during operation of the gas turbine engine 10. Contact of the second end portion 126 and the second rim 132 may help block radial leakage of the hot combustion gases 56. Lengths 125 and 127 of the first and second end portions 124 and 126 may be selected to provide sufficient crush stress and clearance for assembly and removal for the interstage seal 40 depending on the selected materials. For example, the lengths 125 and 127 may be between approximately 5 mm to 50 mm, 10 mm to 25 mm, or 15 mm to 20 mm. Each of the lengths 125 and 127 may be between approximately 5 percent to 40 percent, 10 percent to 25 percent, or 15 percent to 20 percent of an overall length 136 of the interstage seal 40.

[0023] In the illustrated embodiment, the interstage seal assembly 41, of which the interstage seal 40 is one segment of the assembly 41, is annularly disposed (in the circumferential direction 54) between the first and second wheels 66 and 74. Thus, the first and second wheels 66 and 74 form annular structures with the interstage seal assembly 41 extending as an annular structure between the first and second wheels 66 and 74. During operation, the first and second wheels 66 and 74 and the interstage seal assembly 41 rotate about a common axis. The interstage seal assembly 41 may include a 360-degree segmented (e.g., 2 to 100 segments) circular structure that attaches to adjacent first and second wheels 66 and 74 to form a wall that thermally isolates an interstage volume or wheel cavity 134 that forms an air-cooling chamber. FIGS. 4-6 illustrated various steps that may be performed during installation of the interstage seal 40. Removal of the interstage seal 40 may be accomplished by performing these steps in reverse. Starting with the first step, FIG. 4 illustrates a partial cross sectional side view of the interstage seal 40 being pivoted between the first and second turbine stages 62 and 64. As shown in FIG. 4, the second blade segments 76 have been removed to facilitate the installation of the interstage seal 40. The first blade segments 68 remain in place in the first stage 62. In other embodiments, the interstage seal 40 may be installed by removing the first blade segments 68, with the second blade segments 76 remaining in place in the second stage 64. Thus, installation of the interstage seal 40 may not involve removal of both the first and second blade segments 68 and 76, thereby substantially simplifying the installation or removal of the interstage seal 40. Moreover, as shown in FIG. 4, the first and second wheels 66 and 74 remain in place during installation (and removal) of the interstage seal 40, thereby substantially simplifying the installation or removal of the interstage seal 40. During installation of the interstage seal 40, the second end portion 126 is positioned, or hooked, under the second rim 132. Specifically, a corner 148 formed between the inclined support rib 88 and the inner bridge portion 84 is placed adjacent to, in an overlapping relationship with, the second wheel rim 132. In other words, an overlap 149 of the corner 148 and the second rim 132 exists in the axial 50 and radial 52 directions. As a result, the interstage seal 40 may be pivoted, or rotated, about the corner 148 in the direction of the arrow 150 toward the axial axis 50. As shown in FIG. 4, an outer edge 151 of the first end portion 124 may follow an arc 152 as the interstage seal 40 is moved in the direction 150. Thus, the overlap 149 provides sufficient clearance to enable the outer edge 151 to clear the first wheel rim 128 as the interstage seal 40 pivots toward the axial axis 50. At the beginning of the installation process, the inclined support rib 88 may be substantially parallel to a face 153 of the second wheel post portion 78. The configuration of the inclined support rib 88 and the recessed portion 90 enables the overlap 149 to be greater, thereby enabling the outer edge 151 to clear the first wheel rim 128 as indicated by the dashed line 152. In addition, the inclined support portion 90 may overlap a portion 154 of the second wheel support post 78 when viewed cross-sectionally. As described in detail below, two or more inclined support portions 90 may surround the second wheel support post 78. As the installation of the interstage seal 40 proceeds, the first end portion 124 rotates toward the first wheel rim 128 and the inclined support rib 88 moves away from the second wheel support post 78.

[0024] FIG. 5 is a partial cross-sectional side view of the interstage seal 40 being moved along the axial axis 50. As shown in FIG. 5, the interstage seal 40 has been rotated such that the first end portion 124 may be moved under the first wheel rim 128 as indicated by arrow 170. In addition, the second end portion 126 may remain overlapping with the second wheel rim 132. Specifically, the corner 148 may be adjacent to the second wheel rim 132. The overlap 149 enables the first end portion 124 to move under the first wheel rim 128 while the overlap 149 is maintained between the second end portion 126 and the second wheel rim 132. As the interstage seal 40 is moved in the direction of arrow 170, the axial stop 130 may contact the first wheel rim 128 to block further axial movement 50 of the interstage seal 40. In addition, the corner 148 moves at least partially away from the second wheel rim 132.

[0025] FIG. 6 is a partial cross-sectional side view of the interstage seal 40 illustrating the completion of the installation process. As shown in FIG. 6, the axial stop 130 is adjacent to the first wheel rim 128, and the corner 148 has moved away from the second wheel rim 132. The first end portion 124 remains axially 50 overlapped with the first wheel rim 128 and the second end portion 126 remains axially 50 overlapped with the second wheel rim 132. In addition, the second blade segment 76 may be moved in the direction of arrow 180 toward the interstage seal 40. Specifically, the second support feature 120 may be engaged with the first support feature 118. Once the first and second support features 118 and 120 are engaged, the interstage seal 40 may be blocked from moving toward or away from the axial axis 50 of the gas turbine engine 10. Thus, the interstage seal 40 may be self-supporting throughout the rest of the installation process without having to hold or restrain the interstage seal 40 from moving. Finally, the labyrinth seal 100 may be moved in the direction of arrow 182 toward the interstage seal 40. In certain embodiments, the labyrinth seal 100 may be coupled to the case of the gas turbine engine 10 and thus, the labyrinth seal 100 may be installed when the case of the gas turbine engine 10 is mounted. After the gas turbine engine 10 starts, the centrifugal seals 110 (if used) may move outward to radially seal the gaps between the interstage seal 40 and the first and second turbine stages 62 and 64.

[0026] FIG. 7 is a front perspective view of an embodiment of one interstage seal segment 40 of the interstage seal assembly 41. As shown in FIG. 7, the axial stop 130 includes a first face 200 and a second face 202. The first face 200 may face toward the first turbine stage 62 and the second face 202 may face toward the interstage seal 40. In certain embodiments, the first face 200 may be generally flat to correspond to the first wheel 66. In addition, the second face 202 may be curved to provide a greater support for attachment of the axial stop 130 to the first end portion 124. In other embodiments, the second face 202 may be flat. In addition, the axial stop 130 may have a width 204 that is approximately the same as a width of the interstage seal segment 40. Further, the axial stop 130 may be defined by a height 206, which may be selected to provide sufficient surface area for the axial stop 130 to help block undesired axial movement of the interstage seal 40.

[0027] FIG. 8 is a rear perspective view of an embodiment of one interstage seal segment 40 of the interstage seal assembly 41. As shown in FIG. 8, the interstage seal 40 includes two inclined support portions 90. Specifically, the two interstage support portions 90 are located on outer sides 218 of the interstage seal 40. Thus, an interstage support portion gap 220 exists between the two inclined support portions 90. During installation or removal of the interstage seal 40, the second wheel support post 78 may fit into the interstage support portion gap 220. Thus, the interstage seal 40 may be pivoted about the second wheel support post 78 during installation or removal of the interstage seal 40. By pivoting about the second wheel support post 78, the second stage 64 may remain in place during maintenance of the interstage seal 40. In the illustrated embodiment, each of the inclined support portions 90 is defined by a width 222. Together, the widths 222 of the inclined support portions 90 and the interstage support portion gap 220 may be approximately the same as the width 204 of the interstage seal 40. In addition, the inclined support portions 90 may be defmed by a height 224. As shown in FIG. 8, the height 224 is less than a height of the interstage seal 40. In other embodiments, the height 224 may be smaller or greater depending on the amount of incline of the inclined support rib 88 and the support desired for the outer bridge portion 82. For example, if the angle 94 is smaller, the inclined support portions 90 may have greater heights 224 to provide additional support for the outer bridge portion 82. In other embodiments, only one inclined support portion 90 may be provided near the center of the interstage seal 40. In further embodiments, the interstage seal 40 may include more than two inclined support portions 90. Some embodiments may omit the inclined support portion 90.

[0028] FIG. 9 is a front view of three adjacent interstage seals 40 of the segmented interstage seal assembly 41. The assembly 41 may include a plurality of interstage seals 40, such as 2 to 100 seals 40, disposed adjacent to one another to form a complete 360-degree ring about the longitudinal axis 32 of the gas turbine engine 10. As shown in FIG. 9, each of the interstage seals 40 is arcuate in the circumferential direction 54. The assembly 41 of the interstage seals 40 may include outer seals 240 and inner seals 242. Axial slots 246 may be formed in the outer and inner bridge portions 82 and 84 to accommodate the outer and inner axial seals 240 and 242. In other words, the outer and inner seals 240 and 242 extend in the axial direction 50 along the axial slots 246. Outer and inner seals 240 and 242 may be installed between each of the interstage seals 40 of the interstage seal assembly 41 as discussed above. The outer and inner axial seals 240 and 242 may help to block radial leakage of the hot combustible gases 56.

[0029] FIG. 10 is a side view of an embodiment of the interstage seal 40. As shown in FIG. 10, the outer and inner seals 240 and 242 are disposed in axial slots 246 that run along the outer and inner bridge portions 82 and 84. In addition, the interstage seal 40 may include one or more cooling passages 260 to enable the cooling fluid to flow toward the first and second blade segments 68 and 76. For example, in certain embodiments, the cooling passages 260 may enable the cooling fluid to flow from the interstage volume 134 toward the first and second blade segments 68 and 76. In other embodiments, the cooling fluid may flow from a casing structure of the gas turbine engine 10 into the first and second blade segments 68 and 76 through the cooling passages 260. The cooling passages 260 may be formed in the outer and inner bridge portions 82 and 84, the intermediate support beams 86, and/or the inclined support rib 88. The cooling passages 260 formed in the outer and inner bridge portions 82 and 84 may enable the cooling fluid to flow from the interstage volume 134 or the casing structure to the first and second blade segments 68 and 76. The cooling passages 260 formed in the intermediate support beams 86 and/or the inclined support rib 88 may enable the cooling fluid to flow between the recessed portions 98 and the seal cavities 96.

[0030] This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal language of the claims.

Claims

1. A system, comprising:

a multi-stage turbine (22), comprising:

a first turbine stage (62) comprising a first wheel (66) having a plurality of first blade segments (68) spaced circumferentially about the first wheel (66);

a second turbine stage (64) comprising a second wheel (74) having a plurality of second blade segments (76) spaced circumferentially about the second wheel (74); and

an interstage seal (40) extending axially between the first (62) and second (64) turbine stages, wherein the interstage seal (40) is configured to be installed or removed while the first (66) and second (74) wheels remain in place in the respective first (62) and second (64) turbine stages.

2. The system of claim 1, wherein each of the plurality of first blade segments (68) is coupled to the first wheel (66) using a plurality of first mounts (70), and each of the plurality of second blade segments (76) is coupled to the second wheel (74) using a plurality of second mounts (78).

3. The system of claim 2, wherein each first mount (70) comprises a first slot (72) in the first wheel (66) and a first tab (73) in one of the first plurality of blade segments (68), and each second mount (78) compromises a second slot (80) in the second wheel (74) and a second tab (81) in one of the second plurality of blade segments (76).

4. The system of any of claims 1 to 3, wherein the interstage seal (40) is configured to pivot toward an axial axis (50) of the multi-stage turbine (22) during installation of the interstage seal (40), and the interstage seal (40) is configured to pivot away from the axial axis (50) of the multi-stage turbine (22) during removal of the interstage seal (40).

5. The system of any of claims 1 to 4, wherein the interstage seal (40) comprises an inclined support rib (88) at an angle (94) from an inner bridge portion (84) of the interstage seal (40), wherein the inclined support rib (88) enables the interstage seal (40) to pivot toward and away from an axial axis (50) of the multi-stage turbine (22).

6. The system of claim 5, wherein the interstage seal (40) comprises a second recessed portion (98) adjacent to the inclined support rib (88), the second recessed portion (98) is configured to receive a second portion (132) of the second wheel (74) to enable pivotal motion (150) of the interstage seal (40) toward and away from the axial axis (50), and a first recessed portion (98) of the interstage seal (40) is configured to receive a first portion (128) of the first wheel (66) while the interstage seal (40) is moved along the axial axis (50) toward the first turbine stage (62).

7. The system of claim 5 or 6, wherein the angle (94) is less than approximately 75 degrees.

8. The system of any of claims 1 to 4, wherein the interstage seal (40) comprises a first support feature (118) configured to engage with a second support feature (120) disposed on one or more of the plurality of first blade segments (68) or the plurality of second blade segments (76) to block radial movement (52) of the interstage seal (40) toward an axial axis (50) of the multi-stage turbine (22) during installation or removal of the interstage seal (40), and to block circumferential movement (54) of the interstage seal (40) about the axial axis (50) during operation of the multi-stage turbine (22).

9. The system of claim 8, wherein the first support feature (118) comprises a slot and the second support feature (120) comprises a tab.

10. The system of any of claims 1 to 4, wherein the interstage seal (40) comprises an axial end portion (124, 126) configured to engage a wheel rim (128, 132) of the first wheel (66) or the second wheel (74) in a radial direction (52) during operation of the multi-stage turbine (22).

11. The system of any of claims 1 to 4, wherein the interstage seal (40) comprises a centrifugal seal (110) configured to move toward the first turbine stage (62) or the second turbine stage (64) to block radial leakage when radial centrifugal forces are generated during operation of the multi-stage turbine (22).

12. The system of any of claims 1 to 4, wherein the interstage seal (40) comprises one or more seal teeth (104) configured to block interstage axial leakage between the first turbine stage (62) and the second turbine stage (64).

13. The system of any of claims 1 to 4, wherein the interstage seal (40 comprises one or more cooling passages (260) configured to direct a cooling fluid flow toward the plurality of first blade segments (68) or the plurality of second blade segments (76).

14. The system of any of claims 1 to 4, comprising an interstage seal assembly (41) disposed between the first (62) and second (64) turbine stages, wherein the interstage seal assembly (41) comprises a plurality of interstage seals (40).

15. A method, comprising:

positioning a first recessed portion (98) of an interstage seal (40) about a first wheel rim (128) of a turbomachine (22);

pivoting a second recessed portion (98) of the interstage seal (40) toward an axial axis (50) of the turbomachine (22); and

moving the interstage seal (40) along the axial axis (50) toward a second wheel rim (132) of the turbomachine (22) to position the second recessed portion (58) about the second wheel rim (132).

16. The method of claim 15, comprising engaging a first support feature (118) disposed near the first recessed portion (98) of the interstage seal (40) with a second support feature (120) disposed on a portion of a blade segment (68) coupled to the first wheel rim (128) to block radial movement (52) of the interstage seal (40) toward the axial axis (50) of the turbomachine (22) during installation or removal of the interstage seal (40), and to block circumferential movement (54) of the interstage seal (40) about the axial axis (50) during operation of the turbomachine (22).

Drawing