The present disclosure relates to seals and, more specifically, to seals for gas turbine engines.
A gas turbine engine typically includes a fan section, a compressor section, a combustor section, and a turbine section. Air moves into the engine through the fan section. Airfoil arrays in the compressor section rotate to compress the air, which is then mixed with fuel and combusted in the combustor section. The products of combustion are expanded to rotatably drive airfoil arrays in the turbine section. Rotating the airfoil arrays in the turbine section drives rotation of the fan and compressor sections. The compressor section and turbine section each have multiple stages of blades that rotate about a central axis and multiple stages of vanes that are stationary relative to the central axis.
Gas turbine engines may have various gas-flow streams that may be kept separate from one another. The gas-flow streams may be separated by various components. The internal engine environment is exposed to temperature and pressure extremes and other harsh environmental conditions, which may affect the integrity of the components separating different gas-flow streams. Higher pressure gaspath air may tend to leak between airfoil arrays into lower pressure gaspaths. Seals may be included between the components to minimize leakage. Seal degradation may contribute to loss of performance and/or efficiency of the gas turbine engine.
A seal for gas turbine engine is described herein, in accordance with various embodiments. A seal may comprise a first ply having an annular shape and a second ply including a second ply having a circumferentially discontinuous annular shape including an axial split. The second ply may be positioned directly adjacent to at least one of an inner diameter surface or an outer diameter surface of the first ply.
In various embodiments, the first ply and the second ply comprise a “W” seal. The second ply may be nested within the first ply with an inner diameter surface of the second ply in contact with the surface of the first ply. A third ply may be positioned adjacent to at least one of the first ply or the second ply. The third ply may have a circumferentially discontinuous annular shape. The first ply may comprise a first sheet metal and the second ply may comprise a second sheet metal. The first ply may be comprised of a first material and the second ply may be comprised of a second material. At least one of the first material or the second material may comprise a nickel-based alloy. The first material and the second material may comprise the same material.
An engine section of a gas turbine engine is also provided. The engine section may comprise a first engine component and a second engine component disposed adjacent to the first engine component. A seal disposed between the first engine component and the second engine component. The seal may comprise a first ply having an annular shape. The seal may further comprise a second ply having circumferentially discontinuous annular shape including an axial split. The second ply may be positioned directly adjacent to at least one of an inner diameter surface or an outer diameter surface of the first ply.
In various embodiments, the first ply and the second ply comprise a “W” seal. The second ply may be nested within the first ply with an inner diameter surface of the second ply in contact with the surface of the first ply. The second ply may be mechanically loaded against the surface of the first ply. The first ply may be comprised of a first material and the second ply may be comprised of a second material. The first material and the second material may comprise the same material. At least one of the first material or the second material may comprise a nickel-based alloy. The second material may have a greater wear resistance than the first material. The first ply may comprise a first sheet metal and the second ply may comprise a second sheet metal.
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 turbine section. The engine section may comprise a blade outer air seal and a vane adjacent to the blade outer air seal. A seal may be disposed between the blade outer air seal and the vane. The seal may comprise a first ply having an annular shape. The seal may further comprise a second ply having circumferentially discontinuous annular shape including an axial split. The second ply may be positioned directly adjacent to at least one of an inner diameter surface or an outer diameter surface of the first ply.
In various embodiments, the first ply and the second ply comprise a “W” seal. The second ply may be nested within the first ply with an inner diameter surface of the second ply in contact with the surface of the first ply. The second ply may be mechanically loaded against the surface of the first ply. The first ply may be comprised of a first material and the second ply may be comprised of a second material. At least one of the first material or the second material may comprise a nickel-based alloy.
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 36 or engine case via several bearing systems 38, 38-1, and 38-2. Engine central longitudinal axis A-A′ is oriented in the z direction on the provided xyz axis. 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 44 and a low pressure turbine 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 static 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 gas 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, or may comprise any gas turbine engine as desired.
Referring now to
With reference to
In various embodiments, engine section 80 may include stationary annular fluid seals, referred to as a blade outer air seal (BOAS) assembly, circumscribing the blades 70 and defining a radially outward boundary of core airflow path C. A BOAS assembly may include one or more BOAS segments 90 circumferentially arranged to form an annular ring about engine central longitudinal axis A-A′ radially outward of blades 70. BOAS segments 90 may be disposed axially adjacent with respect to vanes 72. During engine operation, vanes 72 and BOAS segments 90 are subjected to different thermal loads and environmental conditions. Cooling air may be provided to BOAS segment 90 and vanes 72 to enable operation of the turbine during exposure to hot combustion gasses produced within the combustion area, as described above. Pressurized air may be diverted from other engine sections as a secondary airflow S and used to cool components within engine section 80. A BOAS segment 90 and vanes 72 may define an inner diameter boundary of secondary airflow S, and engine case structure 36 may define an outer diameter boundary of secondary airflow S. Secondary airflow S may have a higher pressure than core airflow C. A difference in pressure of secondary airflow S may typically be about 100 to 200 pounds per square inch (psi), or 689 to 1379 kilopascals (kPa), greater than core airflow C, producing a tendency for secondary airflow S to leak into core airflow C. One or more seals 100 may therefore be incorporated between stages of BOAS segment 90 and vanes 72 to reduce leakage of cooling air from secondary airflow S into core airflow C. Seal 100 may comprise an annular seal disposed between BOAS segment 90 and vane 72.
With reference to
In various embodiments, at least a portion of seal 100 may include a continuous annular ring extending circumferentially within cavity 106 around of engine section 80. Cavity 106 may expand and contract (axially and/or radially) in response to the thermal and/or mechanical environment of engine section 80. With a continuous ring portion, seal 100 provides a hoop stiffness, which may reduce the loading of seal 100 against a surface 112 at the inner diameter of seal 100 or against a surface 114 at the outer diameter of seal 100. Seal 100 may be disposed within cavity 106 and may generally be free-floating within cavity 106. Seal 100 may be a “W” seal (e.g. a seal having a “W”-shaped cross-section or that form a “W” shape). A “W” seal may be flexible enough axially to expand and seal against the walls of first engine component 102 and a second engine component. Seal 100 may be biased outwardly along engine central longitudinal axis A-A′ to contact aft surface 108 of first engine component 102 and forward surface 110 of second engine component 104. The “W” shape may enable seal 100 to compress and extend as first engine component 102 moves relative to second engine component 104 in response to thermally driven deformations and pressure loads. Thus, seal 100 prevents or reduces leakage airflow passing through cavity 106.
With reference to
With reference to
With reference to
Seal 100 may include a metal, such as titanium, titanium-based alloy, nickel, nickel-based alloy, aluminum, aluminum-based alloy, steel, or stainless steel, or other materials. First ply 120 may include a first material and second ply 130 may include a second material. In various embodiments, first ply 120 may include the same material as second ply 130 and, in various embodiments, first ply 120 may include a different material than second ply 130. For example, first ply 120 may include a material having a lower coefficient of thermal expansion (CTE) relative to a material of second ply 130. The materials of first ply 120 and second ply 130 may be configured according to the operating conditions of the surrounding environment. For example, in engine section 80 (see
With reference to
Returning to
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.
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Number | Date | Country | |
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20180100406 A1 | Apr 2018 | US |