Patent Application
20170306791
The subject matter disclosed herein relates generally to wear liners.
A gas turbine engine typically includes a fan, a compressor, a combustor and a turbine. Stator airfoils are supported on features defined within an inner case. The features typically include grooves or slots that receive flanges known as feet or hooks. The fit of the feet within the grooves of the inner case are typically a clearance fit that accommodates relative thermal growth during operation. The relative movement can cause wear as well as provide an undesired leak path. The tight tolerances make assembly and manufacture difficult while also increasing costs.
In various embodiments, a wear liner is disclosed that comprises a first hook and a second hook connected by a base to provide a generally S-shaped channel. The wear liner includes an outer surface adapted to lie against a first component and an inner surface adapted to lie against a second component. A first engaging end of the first hook is adapted to engage the first component and a second engaging end of the second hook is adapted to engage the second component.
In various embodiments, the first component is a vane foot and the second component is a slot. Also, the first engaging end comprises a curve intersecting a surface of the vane foot and the second engaging end comprises a curve intersecting a narrow portion of a slot. The wear liner may comprise a single sheet of material.
In various embodiments, the wear liner is manufactured by bending a stripped material to form a first hook and a second hook. A base connects the first hook and the second hook, which forms a generally S-shaped channel. The S-shaped channel has an outer surface that is formed to lie against a first component and an inner surface that is adapted to lie against a second component. A first engaging end of the first hook is formed to press against and seal to the first component and a second engaging end of the second hook is formed to press against and seal to the second component.
In various embodiments, the wear liner is installed within a gas turbine engine by inserting the wear liner into an inner diameter of a fan case. One of ordinary skill in the art will appreciate that while the wear liner is described relative to specific implementations (e.g., a fan case), this is for explanation only. The wear liner may be configured with little or no modification, to be suitably implemented within a number of varying modules and components for a gas turbine engine. Moreover, while generally described in the context of a gas turbine engine, the application for the disclosed wear liner is not so limited. The wear liner may be implemented in any application having a foot and hook configuration similar to what is disclosed herein.
The wear liner includes a first hook and a second hook that are connected by a base to provide a generally S-shaped channel. An outer surface of the wear liner is pressed against a first component and an inner surface is pressed against a second component. A first engaging end of the first hook engages the first component and a second engaging end of the second hook engages the second component.
In various embodiments, the wear liner is inserted as a 360° coil. The wear liner may also comprises a first 180° section and a second 180° section that are inserted end-to-end to form a 360° coil. The first engaging end of the first hook is formed to include a curve intersecting a surface of a vane foot and the second engaging end is formed to include a curve intersecting a narrow portion of a slot. The wear liner may be installed as a single sheet of material.
In various embodiments, the disclosed wear liner comprises a single surface, thereby reducing loose fits as may occur in a dual surface wear liner configuration. The disclosed wear liner may increase tolerance to a stator and case interface over a typical dual surface configuration. In contrast, a dual surface wear liner may significantly increase this tolerance due being configured to wrap around both inner and outer diameters of a case groove.
The accompanying drawings are included to provide a further understanding of the present disclosure and are incorporated in, and constitute a part of, this specification, illustrate various embodiments, and together with the description, serve to explain the principles of the disclosure.
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.
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.
For example, in the context of the present disclosure, systems and methods may find particular use in connection with a half wear liner that protects the inside surface of a fan case. However, various aspects of the disclosed embodiments may be adapted for optimized performance with a variety of turbine assemblies. As such, numerous applications of the present disclosure may be realized.
An X-Y-Z coordinate system is shown in
As used herein, terms such as “under”, “below”, “on-top”, “above”, etc., may be used in describing relative position along the axis, with on top and above reflecting positive Z displacement and under and below reflecting negative Z displacement.
Although the disclosed embodiments frequently depict a turbofan gas turbine engine, it should be understood that the concepts described herein are not limited to use with turbofans.
The example gas turbine engine 100 generally includes a low-speed spool 110 and a high-speed spool 112 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 116 via bearing systems 118. It should be understood that bearing systems 118 at various locations may alternatively or additionally be provided.
The low-speed spool 110 generally includes an inner shaft 120 that connects a fan 122 and a low-pressure compressor 124 to a low-pressure turbine 126. The inner shaft 120 drives the fan 122 through a speed change device, such as a geared architecture 128, to drive the fan 122 at a lower speed than the low-speed spool 110. The high-speed spool 112 includes an outer shaft 130 that interconnects a high-pressure compressor 132 and a high-pressure turbine 134. The inner shaft 120 and the outer shaft 130 are concentric and rotate via the bearing systems 118 about the engine central longitudinal axis A.
A combustor 136 is arranged between the high-pressure compressor 132 and the high-pressure turbine 134. In one example, the high-pressure turbine 134 includes at least two stages to provide a double stage type of high-pressure turbine 134. In another example, the high-pressure turbine 134 includes only a single stage. As used herein, a “high-pressure” compressor or turbine experiences a higher pressure than a corresponding “low-pressure” compressor or turbine.
In various embodiments, low-pressure turbine 126 has a pressure ratio that is greater than about 5. The pressure ratio of the example low-pressure turbine 126 is measured prior to an inlet of the low-pressure turbine 126 as related to the pressure measured at the outlet of the low-pressure turbine 126 prior to an exhaust nozzle.
A mid-turbine frame 138 of the engine static structure 116 is arranged generally between the high-pressure turbine 134 and the low-pressure turbine 126. The mid-turbine frame 138 further supports having bearing systems 118 in the turbine 108 as well as setting airflow entering the low-pressure turbine 126.
The core airflow C is compressed by the low-pressure compressor 124 then by the high-pressure compressor 132 mixed with fuel and ignited in the combustor 136 to produce high speed exhaust gases that are then expanded through the high-pressure turbine 134 and low-pressure turbine 126. The mid-turbine frame 138 includes a plurality of stator vane 140, which are in the core airflow path and function as an inlet guide vane for the low-pressure turbine 126. Utilizing the stator vane 140 of the mid-turbine frame 138 as the inlet guide vane for low-pressure turbine 126 decreases the length of the low-pressure turbine 126 without increasing the axial length of the mid-turbine frame 138. Reducing or eliminating the number of vanes in the low-pressure turbine 126 shortens the axial length of the turbine 108. Thus, the compactness of the gas turbine engine 100 is increased and a higher power density may be achieved.
In various embodiments, the gas turbine engine 100 is a high-bypass geared aircraft engine. The gas turbine engine 100 includes a bypass ratio greater than about six (6), with an example embodiment being greater than about ten (10). The example geared architecture 128 is an epicyclical gear train, such as a planetary gear system, star gear system or other known gear system, with a gear reduction ratio of greater than about 2.3.
In one disclosed embodiment, the gas turbine engine 100 includes a bypass ratio greater than about ten (10:1) and the fan diameter is significantly larger than an outer diameter of the low-pressure compressor 124. It should be understood, however, that the above parameters are only exemplary of one embodiment of a gas turbine engine including a geared architecture and that the present disclosure is applicable to other gas turbine engines.
The example gas turbine engine includes a fan 122 that comprises less than about twenty-six (26) fan blades. In various embodiments, the fan 102 includes less than about twenty (20) fan blades. Moreover, the low-pressure turbine 126 includes no more than about six (6) turbine rotors schematically indicated at 114. In various embodiments, the low-pressure turbine 126 includes about three (3) turbine rotors. A ratio between a number of blades of fan 122 and the number of low-pressure turbine rotors is between about 3.3 and about 8.6.
Referring to
In various embodiments, a wear liner 210 is disposed between the vane foot 214 and the inner surfaces of the slot 204. The wear liner 210 is axially installed and provides wear protection for the inner diameter (ID) of the fan case. More particularly, the wear liner 210 provides wear protection to slot 204 along with the vane foot 214, to prevent fretting and/or galling.
Referring to
The contact between a foot surface 224 of the vane foot 214 and the second hook 218 of the wear liner 210 provides a sealing contact between the wear liner 210 and the vane foot 214.
In various embodiments, material properties of the sheet utilized to form the disclosed wear liner 210 are compatible with the temperatures and pressures encountered during operation. The sheet may comprise any material having attributes that may be desired and/or critical for the specific wear liner implementation. The sheet may comprise any suitable metal, ceramic, mineral, or plastic. Further, the surface finish of the wear liner 210 is such that the desired contact seal is formed with the inside-facing side of surface 208 of the slot 204 and the surface of the vane foot 214. Moreover, it is contemplated that the wear liner 210 may include a coating to further inhibit wear and provide the desired sealing properties. Coatings typically used in the aerospace industry are known among those of ordinary skill in the art. Coatings that might be well-suited for various applications may include thermal spray coatings, ceramic coatings, cermet coatings, abradable coatings, etc.
Referring to
Referring to
In various embodiments, the wear liner 500 is constructed through a progressive fabrication process that use press brakes to coin or air-bend stripped material into the shapes described herein, relative to various embodiments. In various embodiments, a stripped material is formed in accordance with the precise bends, angles, and dimensions of a specific gas turbine engine.
In various embodiments, the shape of the disclosed wear liner 500 is formed by way of a series of in-line rollers that progressively bend the stripped material as it is moved through a series of inline rollers. The rollers cause the stripped material to coil to a precise curvature, which simplifies installation and ensures a proper seal between the finished wear liner and a vane foot.
In various embodiments, the wear liner 500 is cut from a coil according to the precise application properties. For example, a continuous coil may be cut to lengths forming a 180° coil (e.g., half of the radius of a fan) or a full 360° coil. It is contemplated that a manufacturing process may produce a wear liner 500 coil having any finished size, radius, or length. For example, the disclosed wear liner 500 may be produced in six (6) 60° segments for portability, wherein the segments are attached end-to-end prior to installation. In various embodiments, multiple segments that together comprise a full 360° wear liner 500, may each be installed individually.
Manufacturing may further include forming a first engaging end (step 615) of the first hook to seal to the first component vane foot. A second engaging end is formed (step 620) of the second hook to seal to the slot.
Benefits, other advantages, and solutions to problems 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, solutions to problems, and any elements that may cause any benefit, advantage, or solution 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. Different cross-hatching is used throughout the figures to denote different parts but not necessarily to denote the same or different materials.
Devices and methods are provided herein. In the detailed description herein, references to “one embodiment”, “an embodiment”, “various embodiments”, 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 herein is to be construed under the provisions of 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.
This invention was made with government support under N00019-02-C-3003 awarded by United States Air Force. The government has certain rights in the invention.
