Patent Application
20170306768
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.
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.
In the drawings:
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.
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
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
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.
Each shroud element 102 as depicted in
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 FC from the dovetail 116 which translates to an upward force F2 at the second radial edge 108 and a downward force F1 from 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
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 F1 and 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
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.
| Number | Date | Country | Kind |
|---|---|---|---|
| P-416301 | Feb 2016 | PL | national |
