KR100473751B1 - Gas turbine engine shroud seals - Google Patents

Abstract

The seal assembly of the turbine blade tip shroud comprises an integrally bent metal layer intersegment seal having axial and vertical elements suitable for fitting into grooves formed in the end walls of each shroud segment. The shroud segment has a platform and radially extending ribs, the seal fits snugly in a groove that provides the seal with axial and radial elements. In addition, a ring seal is provided at an engagement portion between the shroud and the mounting structure, wherein the mounting structure is formed with an annular clearance having an “L” shaped cross section, the ring seal having a “C” shaped shape, and “C” One of the legs of the ruler extends in the axial element of the "L" shaped gap. The legs have a corrugated pattern wherein the peaks of the corrugated pattern contact the opposite radial walls of the axial element of the gap to form a gas seal at the joint.

Description

Gas Turbine Engine Shroud Seals {GAS TURBINE ENGINE SHROUD SEALS}

The present invention relates to a gas turbine engine, and more particularly to a turbine blade shroud seal for reducing the leakage of cooling fluid in the casing surrounding the turbine blades.

Many attempts have been made to improve the sealing performance of the annular shroud surrounding the tip of the turbine blade. The engine casing surrounds the turbine section of the aircraft engine and is configured to deliver cooling gas in the casing to various elements such as the shroud itself. The turbine section of the engine, concentric with the casing, is in the path of very hot gases. The peripheral casing structure is preferably kept at a relatively low temperature.

The annular shroud of the turbine bleed tip is typically divided circumferentially. A seal is needed to prevent gas leakage in the gaps formed between the ends of the shroud segments. U.S. Patent No. 4,767,260 to Clevenger et al., Issued August 30, 1988, and U.S. Patent No. 5,158,430 to Dixon et al., Issued October 27, 1992, are disposed between the platform of the vane assembly. Deformation of feather seals is disclosed. US Patent No. 4,573,866, issued to Sandy, Jr., et al., March 4, 1986, discloses a feather seal in the radial plane between the ends of the blade tip shrouds. U.S. Patent No. 5,318,402 to Bailey et al., June 7, 1994, discloses a circumferential spacer band and a spline or feather seal that couples the shroud and the feather seal together.

In addition, different sealing rings are needed for different shrouds and casing types. For example, in the form of split joints with grooves with axial elements, it is common to use "C" shaped rings. In US Pat. No. 4,537,866, bellows type rings are used for such grooves. Although such ring seals provide good sealing quality, they tend to harden on compression due to their radial cross-section, so they cannot be easily replaced during engine maintenance. Indeed, special tools may be needed to replace such seals.

1 is a partially exploded perspective view showing an embodiment of the present invention;

2 is a front view of the detail shown in FIG.

3 is a partial cross-sectional perspective view of another embodiment of the present invention;

4 is an axial sectional view of a turbine section of a gas turbine engine showing details of the present invention;

FIG. 5 is an enlarged fragmentary sectional view of the detail of FIG. 4 in the same plane as FIG. 4; FIG.

It is an object of the present invention to provide an improved mutual shroud segment seal of the type having an axial element and a radial element.

Another object of the present invention is to provide a simple one-piece sheet seal for sealing the gap formed between the ends of the circumferential shroud segments to reduce the leakage of cooling gas into the hot gas path.

It is a further object of the present invention to provide a ring seal for sealing an annular engagement portion with an annular clearance having an axial element.

It is a further object of the present invention to provide an improved sealing structure of a turbine blade tip shroud assembly that can be easily compressed radially for ease of installation.

Yet another object of the present invention is to provide a ring seal formed from a thin plate.

The structure according to the invention comprises a seal for a turbine blade tip shroud of a gas turbine engine, wherein the shroud is disposed between an axially oriented hot gas path and a concentric cooling gas conduit formed between the shroud and the outer casing, The shroud comprises a plurality of segments arranged circumferentially, each shroud segment having an axial platform and at least one radial rib extending from the platform, the shroud segment Has a wall and a groove, the groove being formed in an end wall portion corresponding to the platform and an end wall portion corresponding to the rib, wherein the seal has a portion of the seal having an axial element and another portion of the seal radially. A one-piece shee formed in the form of a continuous groove in the end wall to have an element t member).

In a more particular embodiment of the invention, the vertical elements of the sheet forming the seal are superimposed so that the vertical elements are elastic members that provide frictional engagement in the grooves formed in the walls of the radial ribs.

In another embodiment of the present invention, a ring seal is provided to the gas turbine engine. The ring seal seals the joint between two annular engine parts when an axial element is formed in the annular gap, the axial gap having a spaced radial wall, the ring seal comprising an undulation Having a portion, wherein one peak portion of the corrugation portion is in contact with one of the separation walls and the other peak portion of the corrugation portion is in contact with the other spaced walls of the gap, such that the gap leaks by gas flowing on both sides of the coupling portion. Sealed against, the corrugation portion corresponding at an angle sufficient to easily compress the seal in the radial direction.

In a more particular embodiment of the invention, a ring seal for a turbine blade tip shroud of a gas turbine engine is provided. The shroud is disposed between an axially directed hot gas path and a concentric cooling gas conduit formed between the shroud and the outer casing, the shroud having an axially extending platform, at least the ribs from which the mounting means on the casing. Extend radially to be coupled to the The annular clearance has an axial element formed between the shroud and the mounting means of the casing, the annular clearance having opposing radially spaced walls formed on the shroud and the mounting means, so that the ring seal has a "C" shaped portion. The leg of the "C" shaped portion extends axially such that the "C" shaped portion has a radial size larger than the axial leg, and the "C" shaped portion is disposed in the radial element of the gap and radiates the gap. It has a radial size smaller than the directional size, and the leg of the ring seal includes a wavy pattern having repeating peaks, the peaks contacting opposite radially spaced walls of the annular clearance to prevent leakage of the cooling gas flow. And a gas seal at the joint between the mounting means and the mounting means, wherein the corrugated pattern is angled enough to easily compress the seal in the radial direction. It corresponds to.

The nature of the invention has been described in general and will now be described with reference to the accompanying drawings showing preferred embodiments of the invention by way of example.

1 and 2 of the accompanying drawings, there is shown a typical turbine blade shroud 12 for use in the engine 10. This shroud 12 comprises a rotor on which the turbine blade 38 is mounted and a plurality of segments 12a ... 12n arranged circumferentially and concentrically (see FIGS. 1 and 4).

The shroud segment 12 of FIG. 1 comprises a platform 14 consisting essentially of an annular flat member having an axial element, and a pair of upstanding ribs 16 each having flanges 20, 22. 18). Ribs 16 and 18 and respective flanges 20 and 22 support the shroud platform 14 as well as serve to define cooling air passages and chambers. In addition, as shown in FIG. 4, the flanges 20, 22 serve to mount the shrouds in the engine casing 32 and the mounting structure 36 as described below. The wall 34 of the engine casing 32 may be provided with a vent opening 33 that allows the introduction of cooling air into the space between the support structure and the casing 32. Ventilation openings 31 are also provided in the support structure 36 to introduce cooling air into the annular channel 35 formed between the shroud 12 and the support structure 36. Finally, a hole 37 in the shroud 12 discharges this cooling air into the hot gas path.

In order to provide mutual sealing of the segments, continuous grooves 26 are provided in the end walls of each shroud segment 12. The continuous groove 26 includes ribs 16, 18 and corresponding axial elements 26b and radial elements 26a, 26c, respectively. The seal 24 consists of a sheet 24 of heat resistant alloy bent to provide the platform 14 and corresponding upstanding elements 28, 30 and horizontal members 25, 27. Of course, the seal can be extruded or molded from any suitable heat resistant material.

Each upright element 28, 30 has legs 28a, 28b, 30a, 30b slightly divergent and wider than each groove as shown in FIG. 2. This shape provides some elasticity to the upright elements 28, 30 such that the upright elements 28, 30 are upright groove segments when the seal 24 is fitted into the grooves 26 at the ends of the shroud 12. The seals fit snugly into the grooves 26 by pressing each other so as to be inserted into the grooves 26a and 26c. Due to the elasticity, the segments can be easily assembled and then separated and the seal 24 can be replaced if repair is needed.

Another embodiment of the invention is shown in FIGS. 3 to 5. The turbine section of engine 10 is shown in part in FIG. 4, and stator vanes 42, 44 other than turbine blade 38 are shown. While the shroud structure is mounted to each stator vane, the shroud 12 surrounds the blade tip 40 of the turbine blade 38.

As shown, the mounting structure 36 in the engine casing 32 supports the shroud 12 as shown in FIGS. 4 and 5. The flange 22 of the shroud 12 may be mounted on the flange 47 of the mounting structure 36, for example. In the same manner, the flange 20 may be supported by the flange of the mounting structure 36. The shape of the mounting structure 36 as well as the ribs 16, 18 allows cooling channels to be formed in and around the shroud. Since the air pressure in the cooling channel 64 may be different from the air pressure around the channel 66, these different channels or chambers must be sealed.

A ring seal 46 is provided in the gap or annular recess 48 formed by the groove formed by the flange 47. In this case. When the flange 22 is inserted, the recess 48 has an "L" shape with an axial element 48a and a radial element 48b. The ring seal 46 has a "C" shaped portion 52 having a radial size that is smaller than the recess 48 in the region of the radial element 48b. The “C” shaped portion 52 includes legs 54 and 56. The extension of leg 54 has a corresponding corrugated pattern at a flat angle to provide a corrugated section 58 having peaks 60, 62. The corresponding angles by the arms 58a and 58b are obtuse angles. The angle may be acute, but should be large enough that the seal can be easily compressed in the radial direction.

The shape of the corrugated section 58 is such that the radial magnitude between the peaks 60, 62 is slightly larger than the radial magnitude between the walls 36a, 22a of the recess 48. The ring seal 46 is made of a heat resistant material or alloy having an intrinsic elastic modulus suitable for maintaining spring-fitted continual contact with the opposing walls 22a, 36b. Thus, the corrugated section 58 is spring fitted into the recess 48 such that the peaks 60, 62 are in continuous contact with the opposing walls 22a, 36a.

Thus, due to the shape of the ring seal 46, the ring seal can easily fit into the recess 48 before the shroud flange 22 is inserted therein during its assembly. The ring seal 46 may consist of an endless ring to be inserted into the recess 48.

The flange 22 may be formed obliquely as shown to be able to be inserted against a spring seal 46 pre-installed in the recess 48. The radial resilience of the corrugations 58 will allow for easy exchange of the ring seal 46 as well as easy insertion of the flange 22 when removing it from the recess 48 during the retention of the shroud.

Thus, the seal and ring seal 46 can cooperate with the shroud 12 for ease of assembly of the seal and shroud segment, thereby making the maintenance of the shroud 12 very easy.

Claims (11)

A turbine blade tip shroud 12 disposed between the axial hot gas path and the concentric cooling gas conduit, wherein the concentric cooling gas conduit is formed between the shroud 12 and the outer casing 32, and The shroud comprises a plurality of segments 12a... 12n arranged circumferentially, wherein each shroud segment 12a .. 12n has at least an extension from the axial platform 14 and the platform 14. With one radial rib 16, 18, the shroud segments 12a... 12n have an end wall and a groove 26, each groove corresponding to the platform 14. A turbine blade tip shroud seal 24 of a gas turbine engine that is formed continuously in a wall and end wall portions corresponding to the ribs 16, 18,        The seal 24 is in the end wall such that a portion of the seal 24 has an axial element 25, 27 and another portion of the seal 24 has a radial element 28, 30. A turbine blade tip shroud seal comprising an integral seat member formed in the form of a continuous groove 26.       2. The sheet member according to claim 1, wherein the sheet member forming the seal 24 is made of a heat resistant material formed in the form of upright ribs 16, 18, wherein the upright ribs are formed at the ends of the shroud segments 12a ... 12n. Turbine blade tip shroud seal, characterized in that it has a pair of diverging legs (28, 30) that can be elastically compressed to fit snugly in the groove (26) in the subwall. 3. The turbine blade tip shroud seal of claim 2, wherein the heat resistant material is a heat resistant alloy that is bent into a desired shape. 3. The shroud (12) according to claim 2, wherein the shroud (12) has a pair of spaced parallel ribs (16, 18) standing upright from the platform (14), and the continuous groove (26) has shroud segments (12a ... 12n). While extending from both ribs of the sheet member, the sheet member forming the seal 24 comprises a pair of spaced upright portions 28, 30 that are parallel and substantially the same, the pair of upright portions of the shroud segment 12a. Turbine blade tip shroud seals, characterized in that they are resilient to fit within the continuous grooves 26 of the end walls of .12n). For sealing the joint between two annular engine parts 12, 36, the joint between the two parts having an annular clearance 48 having an axial element 48a having opposing radially spaced walls. It is a seal of the gas turbine engine to include,        The seal 46 has an axial portion 54 having a corrugated structure, one peak 60 of the corrugated structure being in contact with one of the opposing radially spaced walls 48a of the annular clearance 48. The other peak portion 62 of the corrugated structure contacts the other one 22a of the separation walls of the annular gap 48 to seal the annular gap 48 against leakage by gas flowing on both sides of the engagement portion. Seal of a gas turbine engine.  6. The seal of a gas turbine engine according to claim 5, wherein the seal (46) is made of a heat resistant natural elastic material.  7. The corrugated structure of claim 6, wherein the corrugated structure comprises repeating peaks (60, 62), the repeating peaks contacting opposite radially spaced walls of the annular clearance (48) to provide a gas seal at the joint. And a corresponding corrugated pattern at an angle sufficient to compress the repeating peaks (60, 62) in a radial direction. 8. The engine of claim 7 wherein the engine comprises an engine casing and shroud 12, wherein the shroud 12 is coupled with a mounting structure 47 in the engine casing 32 to form the shroud 12 and the mounting structure. An axially extending platform 22 defining an annular clearance 48 having an axial element 48a between the annular clearances 48, the annular clearance 48 being the shroud 12 and the mounting structure 47. Opposing radially spaced walls (22a, 48a) formed on the cross section, wherein the repeating peaks (60, 62) of the corrugated pattern of the corrugated structure of the axial extension (54) of the seal (46) are shrouds. Contacting the peaks 60, 62 with opposed radial spacer walls 22a, 48a of the annular gap 48 to provide a gas seal at the engagement between the mounting structure 36 and the mounting structure 36. Corresponding gas turbine engine at an angle sufficient to compress radially so that Seal.  9. The annular clearance of claim 8 further comprises a radial element 48b having an opposite radially spaced wall, wherein the seal 46 is adapted to fit within the radial element of the annular clearance 48. An “C” shaped portion 52 having a radial size smaller than the opposing wall of radial element 48b, wherein the axial portion 54 of the seal is an axial element 48a of the annular clearance 48. And a “C” shaped leg extending axially within the seal.   10. The seal of claim 9, wherein the seal (46) is made from a bent sheet of heat resistant alloy.   The seal of claim 10, wherein the seal (46) is an endless ring.