Patent Application
20190112935
The disclosure relates generally to compressor stators in gas turbine engines.
Gas turbine engines typically include a fan section to drive inflowing air, a compressor section to pressurize inflowing air, a combustor section to burn a fuel in the presence of the pressurized air, and a turbine section to extract energy from the resulting combustion gases. The fan section may include a plurality of fan blades coupled to a fan hub. The compressor section and the turbine section typically include a series of alternating rotors (blades) and stators. Gas turbine engine compressor stator vanes tend to twist relative to each other within the stator vane case during engine operation. Twisting tends to reduce gas turbine engine efficiency and compressor stator vane operational life.
In various embodiments the present disclosure provides a gap closing wearliner may comprise a first liner portion disposed within a groove of a compressor vane case, the first liner portion having first surface parallel to an inner face of the groove and a second surface parallel a lower face of the groove, the first liner portion formed to a contour of the groove and configured to receive a tongue of a pedestal having a vane, a gap between the first surface of the first liner portion and the tongue, and a high Coefficient of Thermal Expansion (CTE) material. The gap closing wearliner may further comprise a second liner portion disposed within the groove of the compressor vane case, formed to the contour of the groove, and configured to receive the tongue of the pedestal. In various embodiments, the high CTE material may be disposed between the first liner portion and the second liner portion.
In various embodiments, the high CTE material is disposed behind the first surface of the first liner portion. In various embodiments, the first surface is driven across the gap in response to a change in temperature. In various embodiments, the first surface contacts the tongue in response to the change in temperature. In various embodiments, an interference between the first surface and the tongue limits a change in an angle of attack of the vane at a maximum. In various embodiments, the high CTE material is disposed behind the second surface of the first liner portion. In various embodiments, the first liner portion further comprises a third surface extending over a portion of an upper face of the tongue, wherein the high CTE material is disposed behind the third surface. In various embodiments, the first liner portion comprises at least one of one metal, metal alloy, titanium, titanium alloy, nickel, nickel alloy, plastic, ceramic, or composite. In various embodiments, a portion of the first liner portion is surface treated. In various embodiments, the high CTE material comprises at least one of rubber, silicone rubber, a Room-Temperature-Vulcanization silicone compound, plastic, or polytetrafluoroethylene. In various embodiments, the high CTE material has a CTE between 10−5(1/F°) and 10−3(1/F°).
In various embodiments the present disclosure provides a gas turbine engine may comprise a compressor section configured to compress a gas, a combustor section aft of the compressor section and configured to combust the gas, a turbine section aft of the combustor section configured to extract energy from the gas, and a first gap closing wearliner wherein the first gap closing wearliner comprises a first liner portion disposed within a groove of a compressor vane case, the first liner portion having first surface parallel to an inner face of the groove and a second surface parallel a lower face of the groove, the first liner portion formed to a contour of the groove and configured to receive a tongue of a pedestal having a vane, a gap between the first surface of the first liner portion and the tongue, and a high Coefficient of Thermal Expansion (CTE) material.
In various embodiments, the first gap closing wearliner further comprises a second liner portion disposed within the groove of the compressor vane case, formed to the contour of the groove, and configured to receive the tongue of the pedestal. In various embodiments, the high CTE material is disposed between the first liner portion and the second liner portion. In various embodiments, the first surface is driven across the gap in response to a change in temperature. In various embodiments, an interference between the first surface and the tongue limits a change in an angle of attack of the vane at a maximum. In various embodiments, the gas turbine engine may further comprise a second gap closing wearliner.
In various embodiments, the present disclosure provides a method of manufacturing a gap closing wearliner. The method may comprise forming a first liner portion and a second liner portion to a contour of a tongue of a vane pedestal and a contour of a groove of a compressor vane case, coupling a high CTE material between the first liner portion and the second liner portion, and inserting the first liner portion and the second liner portion into the groove of the compressor vane case.
The forgoing features and elements may be combined in various combinations without exclusivity, unless expressly indicated herein otherwise. These features and elements as well as the operation of the disclosed embodiments will become more apparent in light of the following description and accompanying drawings.
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 disclosures, however, may best be obtained by referring to the detailed description and claims when considered in connection with the drawing figures, wherein like numerals denote like elements.
The detailed description of exemplary embodiments herein makes reference to the accompanying drawings, which show exemplary embodiments by way of illustration and their best mode. While these exemplary embodiments are described in sufficient detail to enable those skilled in the art to practice the disclosures, 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 disclosures. 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.
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 via one or more bearing systems 38 (shown as bearing system 38-1 and bearing system 38-2 in
Low speed spool 30 may generally comprise an inner shaft 40 that interconnects a fan 42, a low pressure (or first) compressor section 44 (also referred to a low pressure compressor) and a low pressure (or first) 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 (“HPC”) 52 (e.g., a second compressor section) and high pressure (or second) turbine section 54. A combustor 56 may be located between HPC 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 HPC 52, mixed and burned with fuel in combustor 56, then expanded over high pressure turbine 54 and low pressure turbine 46. Mid-turbine frame 57 includes airfoils 59 which are in the core airflow path. Low pressure turbine 46, and high pressure turbine 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 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 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 (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.
In various embodiments, the next generation of turbofan engines may be designed for higher efficiency which is associated with higher pressure ratios and higher temperatures in the HPC 52. These higher operating temperatures and pressure ratios may create operating environments that may cause thermal loads that are higher than the thermal loads encountered in conventional turbofan engines, which may shorten the operational life of current components.
In various embodiments, HPC 52 may comprise alternating rows of rotating rotors and stationary stators. Stators may have a cantilevered configuration or a shrouded configuration. More specifically, a stator may comprise a stator vane, a casing support and a hub support. In this regard, a stator vane may be supported along an outer diameter by a casing support and along an inner diameter by a hub support. In contrast, a cantilevered stator may comprise a stator vane that is only retained and/or supported at the casing (e.g., along an outer diameter).
In various embodiments, rotors may be configured to compress and spin a fluid flow. Stators may be configured to receive and straighten the fluid flow. In operation, the fluid flow discharged from the trailing edge of stators may be straightened (e.g., the flow may be directed in a substantially parallel path to the centerline of the engine and/or HPC) to increase and/or improve the efficiency of the engine and, more specifically, to achieve enhanced compression and/or efficiency when the straightened air is compressed and spun by rotor 64.
According to various embodiments and with reference to
With reference now to
In various embodiments, and with reference now to
In various embodiments and with reference now to
In various embodiments and with reference now to
In various embodiments, a liner portion, such as, for example, liner portion 602, first liner portion 402 or second liner portion 404, is comprised of a metal, a metal alloy, a titanium, a titanium alloy, a nickel, a nickel alloy, a plastic, a ceramic, or a composite. In various embodiments, the CTE of a high CTE material may be between 10−6(1/F°) and 10−3(1/F°), or may be between 10−5(1/F°) and 10−3(1/F°), or may be between, 10−4(1/F°) and 10−3(1/F°). In various embodiments, a high CTE material may have a CTE one or two orders of magnitude greater than many metals. In various embodiments the gap, such as axial gap 412, is between 0.0003 inches and 0.09 inches, or between 0.003 inches and 0.05 inches, or between 0.01 inches and 0.03 inches. In various embodiments, a high CTE material may comprise one of a rubber, a synthetic rubber, a silicone rubber, a Room-Temperature-Vulcanization silicone compound, a plastic, or a polytetrafluoroethylene such as, for example, that sold commercially as Teflon®. In various embodiments, the high CTE material may be bonded to the liner portion by an adhesive such as, for example, a two part adhesive. In various embodiments, the liner portion may have a disrupted surface to promote bonding of the high CTE material. In various embodiments, a disrupted surface is one that is etched, sand blasted, or otherwise rough. In various embodiments, a gap closing wear liner such as, for example, gap closing wearliner 400, may be disposed in a forward groove, such as forward groove 316, and configured to receive a forward tongue such as forward tongue 308. In various embodiments, a vane case, such as compressor vane case 300, may comprise more than one gap closing wear liner.
In various embodiments and with reference now to
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 disclosures.
The scope of the disclosures 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.
Systems, methods and apparatus are provided herein. In the detailed description herein, references to “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 embodiment
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.
