CN110005476B

CN110005476B - Overlapping sealing device

Info

Publication number: CN110005476B
Application number: CN201811531048.XA
Authority: CN
Inventors: P.科瑟; S.希文; A.M.法利克-布鲁克斯
Original assignee: Ansaldo Energia Switzerland AG
Current assignee: Ansaldo Energia Switzerland AG
Priority date: 2017-12-15
Filing date: 2018-12-14
Publication date: 2022-11-08
Anticipated expiration: 2038-12-14
Also published as: CN110005476A; EP3498980A1; EP3498980B1

Abstract

A shiplap arrangement of gas turbines having an axis; it comprises the following steps: first and second blades adjacently arranged in a circumferential direction centered on an axis; each blade comprises, in a radial direction, a foot configured to couple the blade to the rotor, a shank portion provided with a downstream closing wall, a platform and an airfoil portion; providing a cooling air cavity in the shank portion between adjacent buckets, externally sealed by axial seals arranged along the gap between adjacent platforms, and downstream sealed by a shiplap joint between adjacent shank closure walls; the overlap coupling comprises a circumferential protruding portion of the shank closure wall of a first blade and a corresponding receiving recess portion in an adjacent shank closure wall; a radial seal arranged along a gap between the projecting portion and the recessed portion of the shiplap coupling, received at one side in a first radial groove realized in the projecting portion, and received at the other side in a second radial groove realized in the recessed portion, to provide a barrier against leakage flow through the shiplap coupling.

Description

Overlapping sealing device

Cross reference to related applications

The present application claims priority from european patent application No. 17207850.3, filed on 12, 15, 2017, the disclosure of which is incorporated by reference.

Technical Field

The present invention relates to a shiplap (shiplay) device for a gas turbine.

Background

Conventional sealing means for sealing the gap, such as rubber seals, polymer seals, adhesive means, or engagement of a projection in a groove (as encountered in particular in the case of sealing between two static elements) are generally known. In a gas turbine, many components are cooled by a cooling air flow to avoid thermal damage. The cooling air flow should be realized with the lowest possible losses in order to maximize the cooling potential.

In order to achieve an efficient sealing between two blade elements in a gas turbine, for example to prevent loss of cooling air due to leakage flows, it is necessary to match the blade elements to one another exactly. However, if it is desired that adjoining components may have some "clearance" (e.g. between two rotor blades in a rotor of a gas turbine, which is necessary due to the strong flow around the blade element caused by the hot working medium during operation), an exact matching of two adjacent shrouds of the blade element is almost impossible, since such a compact construction (as would be necessary for a complete sealing of the radial clearance) may lead to problems (e.g. due to thermal expansion). Also, after installation, the influence of centrifugal forces between the components may be considerable, which may lead to severe wear of conventional sealing devices. For these reasons, so-called "shingles" are used between blades in a gas turbine rotor according to conventional design for sealing leakage flow in the axial direction. The "laps" constitute heat-resistant sealing means, since they are essentially designed from the material of the blade element itself, form an integral constituent part of the blade element, and thus achieve a sealing effect without additional material that may be sensitive to heat or have a different thermal expansion coefficient.

In most cases, the turbine blade has at least one platform element radially on the inside and/or radially on the outside, which platform element, with the blade row installed, abuts on a respectively adjacent platform element of a respectively adjacent blade element by means of two sides of the platform element pointing in the circumferential direction, so that in each case a substantially circumferential gap is formed. On at least one axial edge (in particular the trailing edge), such a turbine blade element may have, on a first side directed in the circumferential direction, a projection extending in the circumferential direction and projecting into the platform element of the adjoining blade element, and on a second side directed in the circumferential direction, a recess accommodating the projection.

The sequential mounting of such blade elements results in each case in the formation of a so-called "lap" between two blade elements. Such overlap is the overlap or engagement area between a shroud element on an axial edge of a vane element and a shroud element on the same axial edge of an adjacent vane element, which is stepped in the flow direction of the working gas. The overlap seals the radially extending gap between the adjacent circumferential sides of the two turbine blades against the escape of cooling air from the auxiliary air circuit, i.e. against leakage flows in the axial direction. Such an overlap is formed by covering the recess on the first side of the adjacent blade element directed in the circumferential direction by means of the projection on the second side of the blade element directed in the circumferential direction, or by the projection engaging in the recess.

Disclosure of Invention

The present invention is therefore based on the object of providing an improved arrangement with an improved sealing effect along a shiplap coupling which reduces the leakage flow from the cooling air cavity.

This object is achieved by a lapping device for a gas turbine having an axis, the lapping device comprising: a first blade and a second blade adjacently arranged along a circumferential direction centered on a gas turbine axis; each blade comprises, in a radial direction, a foot configured for coupling the blade to the rotor, a shank portion provided with a downstream closing wall, a platform and an airfoil portion; between adjacent blades, in the shank portion, a cooling air cavity is provided, which is sealed externally by an axial seal arranged along the gap between adjacent platforms, and downstream by a shiplap joint between adjacent shank closure walls; the overlap joint comprises a circumferential protruding portion of the shank closure wall of a first blade and a corresponding receiving recess portion in an adjacent shank closure wall; a radial seal arranged along a gap between the projecting portion and the recess portion of the lap coupling and received at one side in a first radial groove realized in the projecting portion and at the other side in a second radial groove realized in the recess portion to provide a barrier against leakage flow through the lap coupling. The claimed arrangement ensures that the gap in the shiplap joint can be sealed.

A further object of the invention is to provide a device in which the seal is rigid and fits with clearance in two facing grooves along the overlapping couplings. This reduces the production costs and ensures easy assembly.

It is a further object of the present invention to provide a device in which the seal is hook-shaped defining a convex side directed against leakage flow pressing the seal against the protrusion and a concave side engaging the protrusion within one of the grooves. This ensures a proper and easy positioning and holding of the seal.

It is a further object of the invention to provide a device in which the groove and seal are positioned between two bends along the gap of the shiplap joint. The bend is typically located in a stepped aft region of the platform and corresponds to a circumferential projection of the platform that extends into the platform of an adjacent blade segment. The bend reduces the rate of leakage flow, and the positioning of the seal further reduces the rate of leakage flow.

The invention has been described above with reference to a sealed shiplap arrangement. However, the invention also relates to a blade row comprising such a sealed shiplap arrangement, and in general to a gas turbine for a power plant provided with such a blade row.

It is to be understood that both the foregoing general description and the following detailed description are exemplary and are intended to provide further explanation of the invention as claimed. Other advantages and features of the present invention will become apparent from the following description, the accompanying drawings, and the claims.

The features of the present invention, which are believed to be novel, are set forth with particularity in the appended claims.

Drawings

Further benefits and advantages of the invention will become apparent upon careful reading of the detailed description with appropriate reference to the accompanying drawings.

The invention itself, however, will be best understood by reference to the following detailed description of the invention, which describes exemplary embodiments of the invention when taken in conjunction with the accompanying drawings, wherein:

FIG. 1 is a schematic illustration of a flow diagram of an embodiment of a gas turbine for a power plant;

FIGS. 2 and 3 are schematic views of two adjacent turbine blades of the gas turbine of FIG. 1;

FIG. 4 is a schematic view of a portion of FIG. 3;

FIG. 5 is a schematic enlarged view of the portion of FIG. 4 indicated by reference character V, with an embodiment of a sealing device according to the invention;

FIGS. 6 and 7 are schematic views of the sealing arrangement of FIG. 5 along the axial direction at standard and worst case tolerances, respectively;

fig. 8 and 9 are schematic views of the sealing device of fig. 5 in a radial direction, wherein the seals are arranged in the blade and in the adjacent blade, respectively.

Detailed Description

Technical contents and detailed description of the present invention are described below according to preferred embodiments (not intended to limit the scope of the present invention) in conjunction with the accompanying drawings. Any equivalent variations and modifications made in accordance with the appended claims are intended to be covered by the claims as hereinafter claimed.

The present invention will now be described in detail with reference to the accompanying drawings.

Referring to FIG. 1, a schematic illustration of a flow diagram of an embodiment of a gas turbine for a power plant is shown. Fig. 1 discloses a gas turbine power plant 1 having an axis a and comprising, continuously along a main flow M:

a compressor section 2, which is provided with an inlet 11 for supply of air 10,

a combustor section provided with a burner 3 and a combustion chamber 4, the burner 3 having a plurality of fuel nozzles 6, in the combustion chamber 4 compressed air is mixed with at least one fuel and the mixture is ignited to form a hot gas stream, and

the turbine section 5 in which the hot gas stream expands, thereby producing work on the rotor 7.

Preferably, the rotor 7 is a single piece made up of a plurality of rotor wheels welded together and extends from the compressor 2 to the turbine 5. As is known, a combustor section may be provided with an annular combustor or a plurality of can combustors. The exhaust gases leaving the turbine may be used in a steam generator 8 and the rotor 7 may be connected to a load 9, which load 9 may be a fixed load, such as a generator in a power plant. As is known, the compressor 2 and the turbine 5 include a plurality of stator vanes and a plurality of rotating blades. These rotary blades are connected to the rotor 7 and are arranged in parallel circumferential rows centred on the axis 7.

Referring to fig. 2 and 3, fig. 2 and 3 are schematic views of adjacent blades of the gas turbine of fig. 1, and in particular, fig. 2 and 3 refer to adjacent turbine blades 12. As is known, the turbine blades are in contact with the hot gases and therefore need to be cooled by cooling air (i.e., compressed air). Specifically, FIG. 2 discloses an illustration of the main directions of the gas turbine field. In this illustration, reference numeral 13 refers to an axial direction, which is parallel to the rotor 7, the axis a and substantially parallel to the hot gas flow direction M. The terms downstream and upstream refer to the axial direction 13 along the hot gas flow direction M. Reference numeral 14 denotes a radial direction centered on the axis a; the terms inner/inner and outer/outer refer to the distance from the axis a along the radial direction 14. Reference numeral 15 denotes a circumferential direction centered on the axis a. For clarity, the following fig. 3-9 also disclose the legend to allow for easy identification of the correct location of the represented gas turbine components. In view of the above, FIG. 2 discloses an axially downstream view of two adjacent turbine blades 12 in the same row of blades. The blade 12 may be constructed of a metal, metal alloy, ceramic Matrix Composite (CMC), or other suitable material. Starting from axis a and along radial direction 14, each blade 12 comprises a foot 16 configured to couple with rotor 7, a shank portion 17 provided with a downstream or closing wall 18, a platform 19 and an airfoil 20. Between two adjacent blades, at the shank portion 17, there is a cavity, in particular a cooling air cavity, which is partially disclosed in fig. 8 and 9 with reference number 21. From the shank cavity, the cooling air may enter into a cooling duct realized inside the airfoil 20, which airfoil 20 is in contact with the hot gas flow in operation. The cavity is bounded externally in the radial direction 14 by adjacent edges of the platform 19 and downstream in the axial direction 13 by adjacent edges of the closing wall 18. As is known, to avoid a loss of efficiency of the turbine engine, the cooling air cavity must be sealed both in the radial direction (i.e. to avoid leakages passing to the gaps existing between adjacent platforms 19) and in the axial direction (i.e. to avoid leakages passing to the gaps existing between the closing walls 18). The sealing in the radial direction is performed by axial seals, preferably in the form of axial sealing strips, arranged in axial seats of the facing adjacent edges of the platform 19. Fig. 2 and 3 disclose schematically the axial seal at 22, and fig. 8 and 9 disclose the seat 23 of the axial seal 22 in the platform 19. The sealing in the axial direction is performed by a so-called shiplap joint, indicated by reference numeral 24 in fig. 2 and 3. As is known, the overlapping coupling 24 comprises a radial edge of a closing wall 19 projecting to an adjacent closing wall 19 and a recess realized in such an adjacent closing wall 19 configured for housing the projecting edge. To assemble the row of blades, each blade in the row (except for the last blade) comprises a closing wall in which the first radial edge is provided with a projection along the circumferential direction 15 and on the opposite radial edge with a radial recess suitable to accommodate the projection of the adjacent blade.

Fig. 3 schematically shows a top view of some adjacent blades in the radial direction 14 in order to clarify the overlap coupling. The airfoil 20 includes a leading edge 25 and a trailing edge 26. Fig. 3 discloses a radial gap 27 between two adjacent platforms 19, sealed by the axial seal 22, and a lap joint 24 configured for closing the axial gap between two adjacent closing walls 19. The protruding portion of the overlap coupling 24 is indicated by reference numeral 28 and the corresponding recess is indicated by reference numeral 29.

Reference is made to fig. 4, which is a schematic enlarged view of the overlapping portion of fig. 3. In particular, fig. 4 discloses a seal 30 arranged in the overlap region 24 described previously. The seal 30 is configured to provide a barrier against leakage flow from the shank cavity through the gap existing between the projection portion 28 and the recess portion 29.

Reference is made to fig. 5, which is a schematic enlarged view of the portion of fig. 4, indicated by the reference V and showing a plan view in the radial direction of the overlapping area 24 provided with the seal 30.

According to the embodiment of fig. 5, the seal 30 extends axially from a first groove 31 realized in the protruding portion 28 to a second groove 32 realized in the recessed portion 29. Reference numerals C1 and C2 in fig. 5 denote cooling air flows that pass between two adjacent closing walls (specifically, through gaps existing between the protrusion portions 28 and the recess portions 29). C1 is a leakage flow in the axial direction, and C2 is a leakage flow in the circumferential direction. As disclosed in fig. 5, the seal 30 allows to block the flow C2. Preferably, the seal 30 is a rigid seal that is received in the

grooves

31 and 32 with play or clearance. The first groove 31 and the second groove 32 face each other substantially along the axial direction 13.

Reference is now made to fig. 6 and 7, which are schematic illustrations of the sealing arrangement of fig. 5 in standard and worst case tolerances, respectively.

The play of the seal 30 in the

relative grooves

21 and 32 ensures that the manufacturing tolerances are compensated and at the same time a proper sealing is obtained. Specifically, the clearance is in a direction parallel to the leakage flow C2 (i.e. in the circumferential direction 15) and/or in a direction transverse (preferably, perpendicular) to the leakage flow C2 (i.e. in the axial direction 13). This results in the rigid seal 30 being floatingly mounted in the

seats

31, 32. Moreover, the seal 30 has the shape of a hook defining a convex side 33 directed against the leakage flow C2 and a concave side 34 engaging a projection 35 within one groove 32 (in particular, the groove 32 in the concave portion 29). The coupling between the seal 30 and the projection 35 is a floating coupling due to the clearance of the seal inside the

seats

31, 32. In use, the leakage flow C2 presses the seal 30 against the projection 35, even if centrifugal forces act on the seal 30.

According to the disclosed embodiment, the seal 30 has a substantially "L" cross-section orthogonal to the radial direction 14, defined by a longer arm 36 spanning the gap of the shiplap coupling and a shorter arm 37 housed in the groove 32 in the recess portion 29. The L-shape is easy to manufacture (e.g. by bending). However, it is possible to obtain a similar function with other profiles (e.g. a "T" profile) that can be manufactured (e.g. by extrusion). In use, leakage flow presses the longer arm 36 against the inner surface of the groove 31 in the recess portion 28, even if centrifugal forces act on the seal 30.

Referring to fig. 8 and 9, which are schematic perspective views of the sealing arrangement of fig. 5 in the axial direction, the seals arranged in a blade and in an adjacent blade are shown, respectively. In particular, fig. 8 and 9 disclose two axial perspective views of the seal 30 taken from the shank cavity, respectively, from the outside of the shank cavity. According to this embodiment, the seal 30 and the

grooves

31, 32 extend radially along substantially the entire radial length of the protrusion 28 and the recess 29.

The seal 30 extends outwardly to an axial seal disposed in the seat 23 between adjacent blades.

Finally, the inner end of the groove is open to allow sliding fitting of the seal 30, the seal 30 being held in place at the opposite outer end of the groove. Preferably, the seal 30 is retained by the axial seal 22, and the axial seal 22 is assembled after the seal 30 is assembled.

Although the invention has been explained in relation to preferred embodiments thereof as mentioned above, it is to be understood that many other possible modifications and variations may be made without departing from the scope of the invention. It is therefore contemplated that the appended claims or claims will cover such modifications and variations as fall within the true scope of the invention.

Claims

1. A lapping device for a gas turbine (1) having an axis (A); the overlapping device comprises:

a first blade and a second blade (12) adjacently arranged along a circumferential direction (15) centered on a gas turbine axis (A); each blade (12) comprises, along a radial direction (14), a foot (16) configured for coupling the blade (12) to a rotor (7), a shank portion (17) provided with a downstream closing wall (18), a platform (19) and an airfoil portion (20); -between adjacent blades (12), providing in the shank portion (17) a cooling air cavity (21), said cooling air cavity (21) being externally sealed by an axial seal (22) arranged along the gap between adjacent platforms (19) and downstream sealed by a shiplap joint (24) between adjacent downstream closing walls (18); the overlap coupling (24) comprises a circumferential protruding portion (28) of the downstream closing wall (18) of the first blade (12) and a corresponding receiving recess portion (29) in the adjacent downstream closing wall (18);

a radial seal (30) arranged along a gap between the projecting portion (28) and the recessed portion (29) of the shiplap coupling (24) and received at one side in a first radial groove (31) realized in the projecting portion (28) and at the other side in a second radial groove (32) realized in the recessed portion (29) to provide a barrier against leakage flow through the shiplap coupling.

2. The device according to claim 1, characterized in that said radial seal (30) is rigid and is housed, with clearance, in said first radial groove (31) and in said second radial groove (32), so that said radial seal (30) is mounted floating in said first radial groove (31) and in said second radial groove (32).

3. Device according to claim 2, characterized in that the interspace is in the circumferential direction and/or in the axial direction.

4. The device according to claim 1, wherein the radial seal (30) is hook-shaped defining a convex side (33) directed against the leakage flow and a concave side (34) engaging a protrusion (35) in one of the first radial groove (31) and the second radial groove (32), the leakage flow pressing the radial seal (30) against the protrusion (35).

5. The device according to claim 4, characterized in that the radial seal (30) comprises a longer arm (36) spanning the gap between the projecting portion (28) and the recessed portion (29) of the shiplap coupling (24) and a shorter arm (37) projecting from the longer arm (36) and housed with the projection (35) in the second radial groove (32), the leakage flow pressing the longer arm (36) against the inner surface of the first radial groove (31).

6. The device according to any one of claims 1 to 5, characterized in that the first and second radial grooves (31, 32) and the radial seal (30) are positioned between two bends along the gap between the protruding portion (28) and the recessed portion (29) of the shiplap coupling (24).

7. Device according to any one of claims 1 to 5, characterized in that the inner ends of said first (31) and second (32) radial grooves are open to allow sliding fitting of said radial seal (30).

8. The device according to any one of claims 1 to 5, characterized in that the radial seal (30) is held in position at the outer ends of the first and second radial grooves (31, 32).

9. The arrangement according to any one of claims 1 to 5, characterized in that the radial seal (30) is held in place by the axial seal (22), the axial seal (22) being fitted after the radial seal (30) is fitted.

10. The device of any one of claims 1 to 5, wherein the shingled device is a blade row assembly.

11. A gas turbine for a power plant; the gas turbine (1) has an axis (A) and comprises:

a compressor (2) for compressing air,

a combustor (3, 4) for mixing and combusting compressed air leaving the compressor (2) with at least one fuel,

a turbine (5) for expanding the combusted hot gas stream exiting the combustor (3, 4) and producing work on a rotor (7);

wherein the turbine (5) comprises at least a lapping device according to any one of claims 1 to 10.