Patent Application
20180080477
The present disclosure relates to a gas turbine engine and, more specifically, to a stator vane assembly.
Gas turbine engines typically include a fan section, a compressor section, a combustor section and a turbine section. In general, during operation, air is pressurized in the compressor section and is mixed with fuel and burned in the combustor section to generate hot combustion gases. The hot combustion gases flow through the turbine section, which extracts energy from the hot combustion gases to power the compressor section and other gas turbine engine loads. One or more sections of the gas turbine engine may include a plurality of vane assemblies having vanes interspersed between rotor assemblies that carry the blades of successive stages of the section. Each vane assembly may comprise a plurality of a vanes installed within an engine case to form an annular structure. Installation of vane assemblies is prone to human error.
A vane assembly for a gas turbine engine is described herein, in accordance with various embodiments. A vane assembly may comprise an engine case and an anti-rotation lug coupled to the engine case. The anti-rotation lug may have a forward end and an aft end. A vane cluster may be supported within the engine case. The vane cluster may include an outer shroud with a first slot defined by a forward flange of the outer shroud and with a second slot defined by an aft flange of the outer shroud. The second slot may be configured to receive the aft end of the anti-rotation lug from first direction and wherein the aft flange may be configured to block receipt of the anti-rotation lug from a second direction, which is opposite the first direction.
In various embodiments, a width of the second slot may be less than a width of the anti-rotation lug. The second slot may include a tapered opening. The aft end of the anti-rotation lug may include a tapered geometry and is configured to fit into the tapered opening of the second slot. An aft edge of the aft flange may form an aft wall of the second slot. The vane cluster may be positioned adjacent to a radially inner surface of the engine case. The first direction may be directed from the forward flange toward the aft flange of the outer shroud. The second direction may be directed from the aft flange toward the forward flange of the outer shroud.
A compressor section is also provided. The compressor section may comprise a compressor case and an anti-rotation lug coupled to the compressor case. A vane cluster may include an outer shroud supported within the compressor case. The outer shroud may be configured to receive the anti-rotation lug from first direction. The outer shroud may be configured to block receipt of the anti-rotation lug from a second direction, which is opposite the first direction.
In various embodiments, the outer shroud may comprise a forward flange and an aft flange. The forward flange may define a first slot and wherein the aft flange defines a second slot comprising a tapered opening. The anti-rotation lug may comprise an aft end including a tapered geometry configured to fit into the tapered opening of the second slot. An aft edge of the aft flange may form an aft wall of the second slot. The first direction may be directed from the forward flange toward the aft flange of the outer shroud. The second direction may be directed from the aft flange toward the forward flange of the outer shroud. The vane cluster may further include a plurality of vanes mounted to an inner shroud.
A gas turbine engine is also provided. The gas turbine engine may comprise an engine section comprising at least one of a compressor section or a fan section. The engine section may comprise an engine case and an anti-rotation lug coupled to the engine case. The anti-rotation lug may have a forward end and an aft end. A vane cluster may be supported within the engine case. The vane cluster may include an outer shroud with a first slot defined by a forward flange of the outer shroud and with a second slot defined by an aft flange of the outer shroud. The second slot may be configured to receive the aft end of the anti-rotation lug from first direction. The aft flange may be configured to block receipt of the anti-rotation lug from a second direction, which is opposite the first direction.
In various embodiments, the forward flange may define the first slot. The aft flange may define the second slot comprising a tapered opening. The aft end of the anti-rotation lug may comprise a tapered geometry configured to fit into the tapered opening of the second slot. The first direction may be directed from the forward flange toward the aft flange of the outer shroud. The second direction may be directed from the aft flange toward the forward flange of the outer shroud.
The foregoing features and elements may be combined in various combinations without exclusivity, unless expressly indicated otherwise. These features and elements as well as the operation thereof will become more apparent in light of the following description and the accompanying drawings. It should be understood, however, the following description and drawings are intended to be exemplary in nature and non-limiting.
The subject matter of the present disclosure is particularly pointed out and distinctly claimed in the concluding portion of the specification. A more complete understanding of the present disclosure, however, may best be obtained by referring to the detailed description and claims when considered in connection with the figures, wherein like numerals denote like elements.
All ranges and ratio limits disclosed herein may be combined. It is to be understood that unless specifically stated otherwise, references to “a,” “an,” and/or “the” may include one or more than one and that reference to an item in the singular may also include the item in the plural.
The detailed description of various embodiments herein makes reference to the accompanying drawings, which show various embodiments by way of illustration. While these various embodiments are described in sufficient detail to enable those skilled in the art to practice the disclosure, it should be understood that other embodiments may be realized and that logical, chemical, and mechanical changes may be made without departing from the spirit and scope of the disclosure. Thus, the detailed description herein is presented for purposes of illustration only and not of limitation. For example, the steps recited in any of the method or process descriptions may be executed in any order and are not necessarily limited to the order presented. Furthermore, any reference to singular includes plural embodiments, and any reference to more than one component or step may include a singular embodiment or step. Also, any reference to attached, fixed, connected, or the like may include permanent, removable, temporary, partial, full, and/or any other possible attachment option. Additionally, any reference to without contact (or similar phrases) may also include reduced contact or minimal contact. Cross hatching lines may be used throughout the figures to denote different parts but not necessarily to denote the same or different materials.
As used herein, “aft” refers to the direction associated with the tail (e.g., the back end) of an aircraft, or generally, to the direction of exhaust of the gas turbine engine. As used herein, “forward” refers to the direction associated with the nose (e.g., the front end) of an aircraft, or generally, to the direction of flight or motion.
As used herein, “distal” refers to the direction radially outward, or generally, away from the axis of rotation of a turbine engine. As used herein, “proximal” refers to a direction radially inward, or generally, towards the axis of rotation of a turbine engine.
In various embodiments and with reference to
Gas-turbine engine 20 may generally comprise a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A-A′ relative to an engine static structure or engine case structure 36 via several bearing systems 38, 38-1, and 38-2. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided, including for example, bearing system 38, bearing system 38-1, and bearing system 38-2.
Low speed spool 30 may generally comprise an inner shaft 40 that interconnects a fan 42, a low pressure compressor section 44 and a low pressure turbine section 46. Inner shaft 40 may be connected to fan 42 through a geared architecture 48 that can drive fan 42 at a lower speed than low speed spool 30. Geared architecture 48 may comprise a gear assembly 60 enclosed within a gear housing 62. Gear assembly 60 couples inner shaft 40 to a rotating fan structure. High speed spool 32 may comprise an outer shaft 50 that interconnects a high pressure compressor 52 and high pressure turbine 54. A combustor 56 may be located between high pressure compressor 52 and high pressure turbine 54. A mid-turbine frame 57 of engine case structure 36 may be located generally between high pressure turbine 54 and low pressure turbine 46. Mid-turbine frame 57 may support one or more bearing systems 38 in turbine section 28. Inner shaft 40 and outer shaft 50 may be concentric and rotate via bearing systems 38 about the engine central longitudinal axis A-A′, which is collinear with their longitudinal axes. As used herein, a “high pressure” compressor or turbine experiences a higher pressure than a corresponding “low pressure” compressor or turbine.
The core airflow C may be compressed by low pressure compressor 44 then high pressure compressor 52, mixed and burned with fuel in combustor 56, then expanded over high pressure turbine 54 and low pressure turbine 46. Turbines 46, 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion.
Gas-turbine engine 20 may be, for example, a high-bypass ratio geared aircraft engine. In various embodiments, the bypass ratio of gas-turbine engine 20 may be greater than about six (6). In various embodiments, the bypass ratio of gas-turbine engine 20 may be greater than ten (10). In various embodiments, geared architecture 48 may be an epicyclic gear train, such as a star gear system (sun gear in meshing engagement with a plurality of star gears supported by a carrier and in meshing engagement with a ring gear) or other gear system. Geared architecture 48 may have a gear reduction ratio of greater than about 2.3 and low pressure turbine 46 may have a pressure ratio that is greater than about five (5). In various embodiments, the bypass ratio of gas-turbine engine 20 is greater than about ten (10:1). In various embodiments, the diameter of fan 42 may be significantly larger than that of the low pressure compressor 44, and the low pressure turbine 46 may have a pressure ratio that is greater than about five (5:1). Low pressure turbine 46 pressure ratio may be measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of low pressure turbine 46 prior to an exhaust nozzle. It should be understood, however, that the above parameters are exemplary of various embodiments of a suitable geared architecture engine and that the present disclosure contemplates other turbine engines including direct drive turbofans. A gas turbine engine may comprise an industrial gas turbine (IGT) or a geared aircraft engine, such as a geared turbofan, or non-geared aircraft engine, such as a turbofan, a turboshaft, or may comprise any gas turbine engine as desired.
In various embodiments, an engine section, such as fan section 22, compressor section 24 and/or turbine section 28, may comprise one or more stages or sets of rotating blades and one or more stages or sets of stationary vanes axially interspersed with the associated blade stages but non-rotating about engine central longitudinal axis A-A′. For example, the rotor assemblies may carry a plurality of rotating blades, while each vane assembly 100 may carry a plurality of vanes that extend into the core flow path C. The blades may rotate about engine central longitudinal axis A-A′, while the vanes may remain stationary about engine central longitudinal axis A-A′. The blades may create or extract energy (in the form of pressure) from the core airflow that is communicated through the engine section along the core flow path C. The vanes may direct the core airflow to the blades to either add or extract energy. A plurality of vane assemblies 100 may be disposed throughout the core flow path C to impart desirable flow characteristics on the gas flowing through the core flow path C. Vane assemblies 100 may at least one row of vanes arranged circumferentially about the engine central longitudinal axis A-A′.
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In various embodiments, each vane 110 of vane assembly 100 may be circumferentially retained to the engine at an outer diameter and/or an inner diameter of the vane assembly 100. Vanes 110 may be cantilevered with an attachment point at radially inner end 124 or at radially outer end 122. A radially inner end 124 of vane 110 may couple to an inner shroud 130. Vane assembly 100 may include an inner shroud 130, which may be an inner circumferential fixed structure comprised of one or more segments. In various embodiments, a plurality of vanes 110 may be coupled to a segment of inner shroud 130 to form a vane cluster 112. Radially outer end 122 of vane 110 may couple an outer shroud 132. In various embodiments, vane 110 may be integral with a portion of inner shroud 130 or outer shroud 132. For example, each vane 110 may include a discrete portion of outer shroud 132 integral with the vane 110. Thus, each vane cluster 112 may include a plurality of vanes 110 forming a portion of outer shroud 132, and vanes 110 of the vane cluster 112 may be coupled to a segment of inner shroud 130.
In various embodiments, one or more vane clusters 112 are disposed within and supported by an engine case 140. Engine case 140 may be an aft case of a compressor, i.e., a compressor case, for example. Engine case 140 may have an annular geometry configured to receive a plurality of vane clusters 112 positioned adjacent to a radially inner surface 142 of engine case 140. Vane cluster 112 may be slidably received within engine case 140. An anti-rotation lug 144 may interface with outer shroud 132 to prevent undesired circumferential movement or rotation of vane clusters 112 relative to engine case 140. Anti-rotation lug 144 may fit into a slot 134 in outer shroud 132. Slot 134 and anti-rotation lug 144 may be configured with an installation feature, as described herein.
With reference now to
Anti-rotation lug 144 may be positioned on engine case 140 to extend in the axial direction (z direction) and may terminate in a forward end 146 and an aft end 148. A geometry or shape of forward end 146 of anti-rotation lug 144 may be dissimilar to a geometry or shape of aft end 148 of anti-rotation lug 144. In various embodiments, aft end 148 of anti-rotation lug 144 may include a tapered surface 160, while forward end 146 of anti-rotation lug 144 may comprise a generally square, or non-tapered, geometry.
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Benefits and other advantages have been described herein with regard to specific embodiments. Furthermore, the connecting lines shown in the various figures contained herein are intended to represent exemplary functional relationships and/or physical couplings between the various elements. It should be noted that many alternative or additional functional relationships or physical connections may be present in a practical system. However, the benefits, advantages, and any elements that may cause any benefit or advantage to occur or become more pronounced are not to be construed as critical, required, or essential features or elements of the disclosure. The scope of the disclosure is accordingly to be limited by nothing other than the appended claims, in which reference to an element in the singular is not intended to mean “one and only one” unless explicitly so stated, but rather “one or more.” Moreover, where a phrase similar to “at least one of A, B, or C” is used in the claims, it is intended that the phrase be interpreted to mean that A alone may be present in an embodiment, B alone may be present in an embodiment, C alone may be present in an embodiment, or that any combination of the elements A, B and C may be present in a single embodiment; for example, A and B, A and C, B and C, or A and B and C.
Systems, methods and apparatus are provided herein. In the detailed description herein, references to “various embodiments”, “one embodiment”, “an embodiment”, “an example embodiment”, etc., indicate that the embodiment described may include a particular feature, structure, or characteristic, but every embodiment may not necessarily include the particular feature, structure, or characteristic. Moreover, such phrases are not necessarily referring to the same embodiment. Further, when a particular feature, structure, or characteristic is described in connection with an embodiment, it is submitted that it is within the knowledge of one skilled in the art to affect such feature, structure, or characteristic in connection with other embodiments whether or not explicitly described. After reading the description, it will be apparent to one skilled in the relevant art(s) how to implement the disclosure in alternative embodiments.
Furthermore, no element, component, or method step in the present disclosure is intended to be dedicated to the public regardless of whether the element, component, or method step is explicitly recited in the claims. No claim element is intended to invoke 35 U.S.C. 112(f) unless the element is expressly recited using the phrase “means for.” As used herein, the terms “comprises”, “comprising”, or any other variation thereof, are intended to cover a non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements does not include only those elements but may include other elements not expressly listed or inherent to such process, method, article, or apparatus.
