US20170306768A1

US20170306768A1 - Turbine engine shroud assembly

Info

Publication number: US20170306768A1
Application number: US15/441,597
Authority: US
Inventors: Marek Szrajer; Pawel LOPATA
Original assignee: General Electric Co
Current assignee: General Electric Co
Priority date: 2016-02-29
Filing date: 2017-02-24
Publication date: 2017-10-26
Also published as: CA2958106A1; JP2017198190A; PL416301A1; EP3225794A1; CN107131005A

Abstract

Disclosed herein is an interlocking shroud assembly for a turbine engine with a plurality of radially extending, circumferentially spaced airfoils terminating in a shroud element and having opposing radial sides with first and second interlock elements. Further provided is a method of forming a shroud about a plurality of rotating blades in a turbine engine.

Description

BACKGROUND

Turbine engines, and particularly gas or combustion turbine engines, are rotary engines that extract energy from a flow of pressurized combusted gases passing through the engine onto a multitude of rotating turbine blades.
The rotating turbine blades can be supported by shrouds that are interlocked to form a circumferential casing to the turbine. A Z-shaped interlock is a typical configuration choice for a shrouded blade assembly which requires a pre-twist during manufacturing and assembly. Eliminating the pre-twist while maintaining an interlock configuration would be beneficial for shroud assembly manufacturing.

BRIEF DESCRIPTION

In one aspect, embodiments of relate to a turbine engine comprising a rotor having a plurality of radially extending airfoils spaced circumferentially about the rotor, with the airfoils terminating in a tip, and a shroud assembly circumscribing the airfoils and comprising a shroud element mounted to each tip and having opposing radial sides with first and second interlock elements, wherein the first interlock element of one shroud element mates with a second interlock element of a circumferentially adjacent element to form a plurality of interlocks between adjacent shroud elements about the circumference of the airfoils.
In another aspect, embodiments relate to an interlocking shroud assembly for a turbine engine comprising a plurality of radially extending, circumferentially spaced airfoils terminating in a shroud element and having opposing radial sides with first and second interlock elements, which interlock to form interlocks between circumferentially adjacent airfoils.
In yet another aspect, embodiments relate to a method of forming a shroud about a plurality of rotating blades in a turbine engines comprising forming an interlock between circumferentially adjacent tips of the blades and preloading the interlock.

BRIEF DESCRIPTION OF THE DRAWINGS

In the drawings:

FIG. 1 is a schematic cross-sectional diagram of a turbine engine for an aircraft.

FIG. 2 is an assembled plurality of airfoils.

FIG. 3 is a perspective view of a shroud element.

FIG. 4 is another perspective view of a shroud element.

FIG. 5 is an illustration of a shroud assembly.

FIG. 6 is a cross-sectional view of the shroud assembly of FIG. 5.

FIG. 7 is a cross-sectional view of a second embodiment of the shroud assembly of FIG. 5.

FIG. 8 is a single airfoil assembly.

DETAILED DESCRIPTION

The described embodiments of the present invention are directed to a shroud assembly for an airfoil. For purposes of illustration, the present invention will be described with respect to the turbine for an aircraft turbine engine. It will be understood, however, that the invention is not so limited and may have general applicability within an engine, including compressors, as well as in non-aircraft applications, such as other mobile applications and non-mobile industrial, commercial, and residential applications.
As used herein, the term “forward” or “upstream” refers to moving in a direction toward the engine inlet, or a component being relatively closer to the engine inlet as compared to another component. The term “aft” or “downstream” used in conjunction with “forward” or “upstream” refers to a direction toward the rear or outlet of the engine relative to the engine centerline.
Additionally, as used herein, the terms “radial” or “radially” refer to a dimension extending between a center longitudinal axis of the engine and an outer engine circumference.
All directional references (e.g., radial, axial, proximal, distal, upper, lower, upward, downward, left, right, lateral, front, back, top, bottom, above, below, vertical, horizontal, clockwise, counterclockwise, upstream, downstream, aft, etc.) are only used for identification purposes to aid the reader's understanding of the present invention, and do not create limitations, particularly as to the position, orientation, or use of the invention. Connection references (e.g., attached, coupled, connected, and joined) are to be construed broadly and can include intermediate members between a collection of elements and relative movement between elements unless otherwise indicated. As such, connection references do not necessarily infer that two elements are directly connected and in fixed relation to one another. The exemplary drawings are for purposes of illustration only and the dimensions, positions, order and relative sizes reflected in the drawings attached hereto can vary.
FIG. 1 is a schematic cross-sectional diagram of a turbine engine 10 for an aircraft. The engine 10 has a generally longitudinally extending axis or centerline 12 extending forward 14 to aft 16. The engine 10 includes, in downstream serial flow relationship, a fan section 18 including a fan 20, a compressor section 22 including a booster or low pressure (LP) compressor 24 and a high pressure (HP) compressor 26, a combustion section 28 including a combustor 30, a turbine section 32 including a HP turbine 34, and a LP turbine 36, and an exhaust section 38.
The fan section 18 includes a fan casing 40 surrounding the fan 20. The fan 20 includes a plurality of fan blades 42 disposed radially about the centerline 12. The HP compressor 26, the combustor 30, and the HP turbine 34 form a core 44 of the engine 10, which generates combustion gases. The core 44 is surrounded by core casing 46, which can be coupled with the fan casing 40.
A HP shaft or spool 48 disposed coaxially about the centerline 12 of the engine 10 drivingly connects the HP turbine 34 to the HP compressor 26. A LP shaft or spool 50, which is disposed coaxially about the centerline 12 of the engine 10 within the larger diameter annular HP spool 48, drivingly connects the LP turbine 36 to the LP compressor 24 and fan 20.
The LP compressor 24 and the HP compressor 26 respectively include a plurality of compressor stages 52, 54, in which a set of compressor blades 56, 58 rotate relative to a corresponding set of static compressor vanes 60, 62 (also called a nozzle) to compress or pressurize the stream of fluid passing through the stage. In a single compressor stage 52, 54, multiple compressor blades 56, 58 can be provided in a ring and can extend radially outwardly relative to the centerline 12, from a blade platform to a blade tip, while the corresponding static compressor vanes 60, 62 are positioned upstream of and adjacent to the rotating blades 56, 58. It is noted that the number of blades, vanes, and compressor stages shown in FIG. 1 were selected for illustrative purposes only, and that other numbers are possible.
The blades 56, 58 for a stage of the compressor can be mounted to a disk 59, which is mounted to the corresponding one of the HP and LP spools 48, 50, with each stage having its own disk 59, 61. The vanes 60, 62 for a stage of the compressor can be mounted to the core casing 46 in a circumferential arrangement.
The HP turbine 34 and the LP turbine 36 respectively include a plurality of turbine stages 64, 66, in which a set of turbine blades 68, 70 are rotated relative to a corresponding set of static turbine vanes 72, 74 (also called a nozzle) to extract energy from the stream of fluid passing through the stage. In a single turbine stage 64, 66, multiple turbine vanes 72, 74 can be provided in a ring and can extend radially outwardly relative to the centerline 12, while the corresponding rotating blades 68, 70 are positioned downstream of and adjacent to the static turbine vanes 72, 74 and can also extend radially outwardly relative to the centerline 12, from a blade platform to a blade tip. It is noted that the number of blades, vanes, and turbine stages shown in FIG. 1 were selected for illustrative purposes only, and that other numbers are possible.
The blades 68, 70 for a stage of the turbine can be mounted to a disk 71, which is mounted to the corresponding one of the HP and LP spools 48, 50, with each stage having its own disk 71, 73. The vanes 72, 74 for a stage of the compressor can be mounted to the core casing 46 in a circumferential arrangement.
The portions of the engine 10 mounted to and rotating with either or both of the spools 48, 50 are also referred to individually or collectively as a rotor 53. The stationary portions of the engine 10 including portions mounted to the core casing 46 are also referred to individually or collectively as a stator 63.
In operation, the airflow exiting the fan section 18 is split such that a portion of the airflow is channeled into the LP compressor 24, which then supplies pressurized ambient air 76 to the HP compressor 26, which further pressurizes the ambient air. The pressurized air 76 from the HP compressor 26 is mixed with fuel in the combustor 30 and ignited, thereby generating combustion gases. Some work is extracted from these gases by the HP turbine 34, which drives the HP compressor 26. The combustion gases are discharged into the LP turbine 36, which extracts additional work to drive the LP compressor 24, and the exhaust gas is ultimately discharged from the engine 10 via the exhaust section 38. The driving of the LP turbine 36 drives the LP spool 50 to rotate the fan 20 and the LP compressor 24.
A remaining portion of the airflow 78 bypasses the LP compressor 24 and engine core 44 and exits the engine assembly 10 through a stationary vane row, and more particularly an outlet guide vane assembly 80, comprising a plurality of airfoil guide vanes 82, at the fan exhaust side 84. More specifically, a circumferential row of radially extending airfoil guide vanes 82 are utilized adjacent the fan section 18 to exert some directional control of the airflow 78.
Some of the ambient air supplied by the fan 20 can bypass the engine core 44 and be used for cooling of portions, especially hot portions, of the engine 10, and/or used to cool or power other aspects of the aircraft. In the context of a turbine engine, the hot portions of the engine are normally the combustor 30 and components downstream of the combustor 30, especially the turbine section 32, with the HP turbine 34 being the hottest portion as it is directly downstream of the combustion section 28. Other sources of cooling fluid can be, but is not limited to, fluid discharged from the LP compressor 24 or the HP compressor 26. This fluid can be bleed air 77 which can include air drawn from the LP or HP compressors 24, 26 that bypasses the combustor 30 as cooling sources for the turbine section 32. This is a common engine configuration, not meant to be limiting.
FIG. 2 illustrates a plurality of radially extending circumferentially spaced airfoils, or blades 70, with each blade 70 extending from a root 88 and terminating in a tip (FIG. 3) arranged in a circumferential row and supported by an arcuate inner band 96 and an arcuate outer band 98. The arcuate outer band 98 comprises a shroud assembly 100 made up of separate individual shroud elements 102, having opposing radial sides 107, 109, which together circumscribe the blades 70.
Each shroud element 102 as depicted in FIGS. 3 and 4 is integrally formed with the blade 70 at the tip 90 and comprises a flange 104. The flange 104 includes a first and a second radial edge 106, 108 and a seat 110 that projects circumferentially beyond the first radial edge 106 illustrated in FIG. 3. The seat 110 is formed by the first radial edge 106 and the flange 104 in a portion 112 of the flange 104 that is circumferentially outboard of the first radial edge 106 illustrated in FIG. 4.
FIG. 5 illustrates an airfoil assembly 114 in which the shroud element 102 is integrally formed with the blade 70 wherein the blade terminates in a dovetail 116. The dovetail 116 is formed to mount to the rotor 53. When assembled, the blade 70 is sprung 118 to apply a preload to the interlocks. The blade 70 can be sprung as shown by solid line 118 from a predominantly parallel position 120 with respect to a neutral axis 122 of the airfoil assembly 114 to a bowed position 124 when interlocked. The blade 70 can also be sprung 118 from an initial position of predominantly bowed 126 to a parallel position 128 when interlocked.
Regardless of the initial or final positions of the blade 70, the final position 124, 128 will cause the second radial edge 108 bias outwardly and the seat to bias inwardly. This bias is caused by the compressive force F_Cfrom the dovetail 116 which translates to an upward force F₂at the second radial edge 108 and a downward force F₁from the seat 110.
Circumferentially adjacent shroud elements 102 interlock together forming a plurality of interlocks 130 between adjacent shroud elements 102 to form the shroud assembly 100 as illustrated in FIG. 6. A cross-sectional view of an exemplary embodiment of the shroud assembly 100 of FIG. 6 is shown in FIG. 7 where a first flange 132 having a first interlock element 134 mates with a second flange 136 having a second interlock element 138 where the first flange 132 overlies and abuts the second flange 136.
When the shroud assembly 100 is assembled, the second radial edge 108 of the second flange 136 will bias toward the first flange 132 due to the forces F₁and F₂. This bias enables friction forces to form between the first and second interlock element surfaces that bond each shroud element 102 to the next radially adjacent shroud element 102.
A second embodiment of the shroud assembly is contemplated in FIG. 8. The second embodiment is similar to the first embodiment, therefore, like parts will be identified with like numerals increasing by 100 respectively, with it being understood that the description of the like parts of the first embodiment applies to the additional embodiments, unless otherwise noted.
In the second embodiment a first interlock element 234 formed on a first flange 232 includes an angled seat 210 formed to receive an angled second interlock element 238 formed on a second flange 236. The flange 204 of each shroud element 202 therefore includes an angled seat 210 and an angled portion 240 formed to fit into the angled seat 210. While illustrated as two ramps 242, 244 forming an apex 246, the angled seat 210 and angled portion 240 can be any shape where the first interlock element 234 is formed to receive the second interlock element 238.
A method of forming a shroud assembly, comprising a shroud element integrally formed with a blade, about a plurality of rotating blades in a turbine engines includes forming an interlock between circumferentially adjacent tips of the blades and preloading the interlock. The preloading of the interlock where an interlock element is made to bias towards another interlock element.
The embodiments described herein have benefits regarding production, performance, and damping capability. Prior art for shroud blade assemblies include Z-shaped interlocks. Implementing airfoil and shroud radial bending ensures contact between interlock elements to achieve outer band preload at operating conditions without typical torsional bending used in Z-shaped shroud design which requires a pre-twist. This type of bending also only requires blade balancing for centrifugal forces and improves damping at blade resonant vibrations due to increased contact areas between the interlock element surfaces. This increase in contact area also provides for a reduction in outer flowpath leakages. The simplified shape of design and removing a need for a pre-twist, eases manufacturing.
It should be appreciated that application of the disclosed design is not limited to turbine engines with fan and booster sections, but is applicable to turbojets and turbo engines as well.
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 languages of the claims.

Claims

What is claimed is:

1. A turbine engine comprising:

a rotor having a plurality of radially extending airfoils spaced circumferentially about the rotor, with the airfoils terminating in a tip; and

a shroud assembly circumscribing the airfoils and comprising a shroud element mounted to each tip and having opposing radial sides with first and second interlock elements;

wherein the first interlock element of one shroud element mates with a second interlock element of a circumferentially adjacent element to form a plurality of interlocks between adjacent shroud elements about the circumference of the airfoils.

2. The turbine engine of claim 1, wherein the shroud element is integrally formed with the airfoil.

3. The turbine engine of claim 2, wherein the airfoil is a blade.

4. The turbine engine of claim 3, wherein the blade terminates in a dove tail opposite the tip and the dove tail is mounted to the rotor.

5. The turbine engine of claim 1, wherein the airfoil is sized such that the airfoil is sprung when interlocked with adjacent airfoils to apply a preload to the interlocks.

6. The turbine engine of claim 1, wherein the first interlock comprises a first flange, the second interlock comprises a second flange, and the first flange overlies and abuts the second flange.

7. The turbine engine of claim 6, wherein the first interlock further comprises a seat spaced circumferentially from the first flange and the second flange sits within the seat.

8. The turbine engine of claim 7, wherein the seat is formed by a first radial edge and the first flange.

9. The turbine engine of claim 8, wherein the first flange projects circumferentially beyond the first radial edge.

10. The turbine engine of claim 9, wherein the first flange is circumferentially outboard of the first radial edge.

11. The turbine engine of claim 10, wherein the second flange is a second radial edge of an adjacent shroud element.

12. The turbine engine of claim 11, wherein the airfoil is sized such that the airfoil is sprung when interlocked with adjacent airfoils to apply a preload to the interlocks causing the second radial edge to bias toward the first flange.

13. An interlocking shroud assembly for a turbine engine comprising a plurality of radially extending, circumferentially spaced airfoils terminating in a shroud element and having opposing radial sides with first and second interlock elements, which interlock to form interlocks between circumferentially adjacent airfoils.

14. The interlocking shroud assembly of claim 13, wherein the airfoil is sized such that the airfoil is sprung when interlocked with adjacent airfoils to apply a preload to the interlocks.

15. The interlocking shroud assembly of claim 14, wherein the first interlock comprises a first flange, the second interlock comprises a second flange and the first flange overlies and abuts the second flange.

16. The interlocking shroud assembly of claim 15, wherein the first interlock further comprises a seat spaced circumferentially from the first flange and the second flange sits within the seat.

17. The interlocking shroud assembly of claim 16, wherein the seat is formed by a first radial edge and the first flange.

18. The interlocking shroud assembly of claim 17, wherein the first flange projects circumferentially beyond the first radial edge.

19. The interlocking shroud assembly of claim 18, wherein the first flange is circumferentially outboard of the first radial edge.

20. The interlocking shroud assembly of claim 19, wherein the second flange is a second radial edge of an adjacent shroud element.

21. A method of forming a shroud about a plurality of rotating blades in a turbine engines comprising:

forming an interlock between circumferentially adjacent tips of the blades; and

preloading the interlock.

22. The method of claim 21, wherein forming the interlock comprises forming an interlock on circumferentially opposite sides of the tip of the blade.

23. The method of claim 21, wherein preloading the interlock comprises springing the blade.