Patent Application
20200056483
Exemplary embodiments of the present disclosure relate generally to rotor stacks and, in one embodiment, to a reduced static friction coating for a snap location of a rotor stack.
A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. Air entering the compressor section is compressed and delivered into the combustion section where it is mixed with fuel and ignited to generate a high-energy exhaust gas flow. The high-energy exhaust gas flow expands through the turbine section to drive the compressor and the fan section. The compressor section typically includes low and high pressure compressors, and the turbine section includes low and high pressure turbines.
The gas turbine engine includes a plurality of rotors arranged along an axis of rotation of the gas turbine engine, in both the compressor section and the turbine section. At least some of these rotors are connected to axially adjacent rotors, spacers or other rotating components via an interference fit, also known in the art as a “snap fit.”
The areas surrounding the interference fit and the surfaces forming the interference fit can experience a significant amount of wear and stress. Accordingly, improved materials are desired for a more effective and efficient interference fit.
According to an aspect of the disclosure, a gas turbine engine is provided and includes a first rotating component having a first snap surface and a second rotating component having a second snap surface. The first and second snap surfaces are configured to interlock along an interface and at least one of the first and second snap surfaces comprising a tailored-friction material at the interface.
In accordance with additional or alternative embodiments, the first and second rotating components are adjacent to one another.
In accordance with additional or alternative embodiments, at least one of the first and second rotating components includes a platform or a spacer.
In accordance with additional or alternative embodiments, the tailored-friction material includes a lubricious oxide.
In accordance with additional or alternative embodiments, the lubricious oxide includes at least one of ruthenium (Ru), molybdenum (Mo), tungsten (W) and niobium (Nb).
In accordance with additional or alternative embodiments, a thickness of the tailored-friction material at the interface is less than or equal to 1 micrometer.
In accordance with additional or alternative embodiments, one of the first and second snap surfaces faces radially inwardly at the interface and the other of the first and second snap surfaces faces radially outwardly at the interface.
In accordance with additional or alternative embodiments, one of the first and second snap surfaces includes a radially inwardly facing surface and an axial surface facing in a first axial direction and the other of the first and second snap surfaces comprises a radially outwardly facing surface and an axial surface facing in a second axial direction opposite the first axial direction.
According to another aspect of the disclosure, a gas turbine engine is provided and includes a first rotating component having an aft snap surface at an aft edge thereof, a second rotating component having a forward snap surface at a forward edge thereof and third, fourth, fifth and sixth rotating components respectively having aft and forward snap surfaces at respective aft and respective forward edges thereof. Each aft snap surface of the first, third, fourth, fifth and sixth rotating components are configured to interlock with a corresponding forward snap surface of the third, fourth, fifth, sixth and second rotating components along first, second, third, fourth and fifth interfaces, respectively. At least one of the aft and forward snap surfaces includes a tailored-friction material at at least one of the first, second, third, fourth and fifth interfaces.
In accordance with additional or alternative embodiments, the first and third rotating components are adjacent to one another, the third and fourth rotating components are adjacent to one another, the fourth and fifth rotating components are adjacent to one another, the fifth and sixth rotating components are adjacent to one another and the sixth and second rotating components are adjacent to one another.
In accordance with additional or alternative embodiments, at least one of the first, third, fourth, fifth, sixth and second rotating components includes a platform or a spacer.
In accordance with additional or alternative embodiments, the tailored-friction material includes a lubricious oxide.
In accordance with additional or alternative embodiments, the lubricious oxide includes at least one of ruthenium (Ru), molybdenum (Mo), tungsten (W) and niobium (Nb).
In accordance with additional or alternative embodiments, a thickness of the tailored-friction material at the at least one of the first, second, third, fourth and fifth interfaces is less than or equal to 1 micrometer.
In accordance with additional or alternative embodiments, at one or more of the first, second, third, fourth and fifth interfaces one of the aft and forward snap surfaces faces radially inwardly and the other of the aft and forward snap surfaces faces radially outwardly.
In accordance with additional or alternative embodiments, at one or more of the first, second, third, fourth and fifth interfaces one of the aft and forward snap surfaces comprises a radially inwardly facing surface and an axial surface facing in a first axial direction and the other of the aft and forward snap surfaces comprises a radially outwardly facing surface and an axial surface facing in a second axial direction opposite the first axial direction.
According to another aspect of the disclosure, a method of interlocking aft and forward snap surfaces of first and second rotating components of a gas turbine engine is provided. The method includes energetically applying a precursor material of a lubricious oxide to at least one of the aft and forward snap surfaces and executing a heat treatment of the precursor material to generate the lubricious oxide as a film on the at least one of the aft and forward snap surfaces.
In accordance with additional or alternative embodiments, the energetically applying includes at least one of thermal spraying, atomic layer deposition, chemical deposition and plasma deposition.
In accordance with additional or alternative embodiments, the executing of the heat treatment includes heating the precursor material to encourage oxidization thereof.
In accordance with additional or alternative embodiments, the method further includes thermally adjusting relative sizes of the first and second rotating components until the first and second rotating components are fittable together and thermally releasing at least one of the first and second rotating components to assume an original size thereof.
These and other advantages and features will become more apparent from the following description taken in conjunction with the drawings.
The following descriptions should not be considered limiting in any way. With reference to the accompanying drawings, like elements are numbered alike:
These and other advantages and features will become more apparent from the following description taken in conjunction with the drawings.
A detailed description of one or more embodiments of the disclosed apparatus and method are presented herein by way of exemplification and not limitation with reference to the Figures.
The exemplary gas turbine engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided, and the location of bearing systems 38 may be varied as appropriate to the application.
The low speed spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a low pressure compressor 44 and a low pressure turbine 46. The inner shaft 40 is connected to the fan 42 through a speed change mechanism, which in exemplary gas turbine engine 20 is illustrated as a geared architecture 48 to drive the fan 42 at a lower speed than the low speed spool 30. The high speed spool 32 includes an outer shaft 50 that interconnects a high pressure compressor 52 and high pressure turbine 54. A combustor 56 is arranged in the gas turbine engine 20 between the high pressure compressor 52 and the high pressure turbine 54. The engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46. The engine static structure 36 further supports the bearing systems 38 in the turbine section 28. The inner shaft 40 and the outer shaft 50 are concentric and rotate via bearing systems 38 about the engine central longitudinal axis A which is collinear with their longitudinal axes.
The core airflow is compressed by the low pressure compressor 44 and then the high pressure compressor 52, is mixed and burned with fuel in the combustor 56 and is then expanded over the high pressure turbine 54 and the low pressure turbine 46. The high and low pressure turbines 54 and 46 rotationally drive the low speed spool 30 and the high speed spool 32, respectively, in response to the expansion. It will be appreciated that each of the positions of the fan section 22, compressor section 24, combustor section 26, turbine section 28, and fan drive gear system 48 may be varied. For example, geared architecture 48 may be located aft of the combustor section 26 or even aft of the turbine section 28, and the fan section 22 may be positioned forward or aft of the location of geared architecture 48.
The gas turbine engine 20 in one example is a high-bypass geared aircraft engine. In a further example, the gas turbine engine 20 bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10), the geared architecture 48 is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine 46 has a pressure ratio that is greater than about five. In one disclosed embodiment, the gas turbine engine 20 bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor 44, and the low pressure turbine 46 has a pressure ratio that is greater than about five 5:1. Low pressure turbine 46 pressure ratio is pressure measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle. The geared architecture 48 may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present disclosure is applicable to other gas turbine engines including direct drive turbofans.
A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section 22 of the gas turbine engine 20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet (10,688 meters). The flight condition of 0.8 Mach and 35,000 ft (10,688 meters), with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram °R)/(518.7°R)]0.5. The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second (350.5 m/sec).
Referring now to
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Referring now to
In their respective free, unrestrained states and when they are unjoined, the adjacent component snap surface 80 is larger than the rotor snap surface 78. Thus, in order to accomplish the joining at the snap location R5-R6, the compressor rotor 62 may be heated and/or the adjacent component 82 may be cooled to temporarily enlarge the rotor snap surface 78 relative to the adjacent component snap surface 80. Once the rotor snap surface 78 and the adjacent component snap surface 80 are subsequently joined and returned to ambient temperature, the desired interference fit is achieved between the rotor snap surface 78 and the adjacent component snap surface 80.
Materials currently used at the snap locations R4-R5, R5-R6, R6-R7 and R7-R8 typically include PWA1115, PWA1193, PWA1227 and PWA36280 and the design of the rotor stack interface (i.e., the interference fit between the rotor snap surface 78 and the adjacent component snap surface 80 of
In accordance with embodiments, the rotor stack interface includes surfaces, such as the rotor snap surface 78 and the adjacent component snap surface 80 of
With reference to
The first rotating component 102 may be provided as a rotor stack and has an aft snap surface 1021 at an aft edge thereof. The second rotating component 103 may be provided as a rotor stack and has a forward snap surface 1032 at a forward edge thereof. The third rotating component 104 may be provided as a rotor stack and has aft and forward snap surfaces 1041 and 1042 at aft and forward edges thereof. The fourth rotating component 105 may be provided as a rotor stack and has aft and forward snap surfaces 1051 and 1052 at aft and forward edges thereof. The fifth rotating component 106 may be provided as a rotor stack and has aft and forward snap surfaces 1061 and 1062 at aft and forward edges thereof. The sixth rotating component 107 may be provided as a rotor stack and has aft and forward snap surfaces 1071 and 1072 at aft and forward edges thereof.
The aft snap surface 1021 is configured to interlock with the forward snap surface 1042 along a first interface 110 (i.e., the R4-R5 snap location). The aft snap surface 1041 is configured to interlock with the forward snap surface 1052 along a second interface 120 (i.e., the R5-R6 snap location). The aft snap surface 1051 is configured to interlock with the forward snap surface 1062 along a third interface 130 (i.e., the R6-R7 snap location). The aft snap surface 1061 is configured to interlock with the forward snap surface 1072 along a fourth interface 140 (i.e., the R7-R8 snap location). The aft snap surface 1071 is configured to interlock with the forward snap surface 1032 along a fifth interface 150 (i.e., the R8 snap location).
At least one of the aft snap surfaces 1021, 1041, 1051, 1061 and 1071 or at least one of the forward snap surfaces 1032, 1042, 1052, 1062 and 1072 includes a tailored-friction material film 501 (see
In accordance with embodiments, at one or more of the first, second, third, fourth and fifth interfaces 110, 120, 130, 140 and 150, one of the aft and forward snap surfaces faces radially inwardly (see, e.g., the adjacent component snap surface 80 of
With reference to
As shown in
With reference to
With reference to
Benefits of the features described herein are the provision of a tailored-friction film that offers reduced or tailored high temperature static friction capabilities at snap locations of rotor stack regions. This will significantly reduce stresses by allowing for easier movement between interfacing components due to thermal expansion. In addition, this technology may be applied to other engine interfaces where low static friction is desired.
The term “about” is intended to include the degree of error associated with measurement of the particular quantity based upon the equipment available at the time of filing the application.
The terminology used herein is for the purpose of describing particular embodiments only and is not intended to be limiting of the present disclosure. As used herein, the singular forms “a”, “an” and “the” are intended to include the plural forms as well, unless the context clearly indicates otherwise. It will be further understood that the terms “comprises” and/or “comprising,” when used in this specification, specify the presence of stated features, integers, steps, operations, elements, and/or components, but do not preclude the presence or addition of one or more other features, integers, steps, operations, element components, and/or groups thereof.
While the present disclosure has been described with reference to an exemplary embodiment or embodiments, it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the present disclosure. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the present disclosure without departing from the essential scope thereof. Therefore, it is intended that the present disclosure not be limited to the particular embodiment disclosed as the best mode contemplated for carrying out this present disclosure, but that the present disclosure will include all embodiments falling within the scope of the claims.
