Patent Application
20190316479
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-speed exhaust gas flow. The high-speed 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 high pressure turbine drives the high pressure compressor through an outer shaft to form a high spool, and the low pressure turbine drives the low pressure compressor through an inner shaft to form a low spool. The fan section may also be driven by the low inner shaft. A direct drive gas turbine engine includes a fan section driven by the low spool such that the low pressure compressor, low pressure turbine and fan section rotate at a common speed in a common direction.
A blade outer air seal for a gas turbine engine according to an example of the present disclosure includes a seal arc-segment that defines a gaspath side, a non-gaspath side, leading and trailing ends, and first and second circumferential sides. A portion of the gaspath side has a geometrically segmented coating section. The geometrically segmented coating section has a wall that has an array of cells, and a coating is disposed in the array of cells.
In a further embodiment of any of the foregoing embodiments, a remaining portion of the gaspath side includes a non-segmented coating.
In a further embodiment of any of the foregoing embodiments, the array of cells includes parallel elongated grooves.
In a further embodiment of any of the foregoing embodiments, the array of cells includes sloped grooves.
In a further embodiment of any of the foregoing embodiments, the sloped grooves are circumferentially sloped.
In a further embodiment of any of the foregoing embodiments, the sloped grooves are sloped at an angle of 20° to 45°.
In a further embodiment of any of the foregoing embodiments, the array of cells includes grooves, and the grooves intersect.
In a further embodiment of any of the foregoing embodiments, the grooves intersect at non-perpendicular angles.
In a further embodiment of any of the foregoing embodiments, the array of cells includes cylindrical cells.
In a further embodiment of any of the foregoing embodiments, the array of cells includes polygonal cells.
In a further embodiment of any of the foregoing embodiments, the array of cells includes at least one of parallel elongated grooves, sloped grooves, intersecting grooves, cylindrical cells, or polygonal cells.
In a further embodiment of any of the foregoing embodiments, the geometrically segmented coating section is on the leading end of the seal arc-segment.
In a further embodiment of any of the foregoing embodiments, the geometrically segmented coating is on the first circumferential side of the seal arc-segment, and the second circumferential side excludes the geometrically segmented coating.
In a further embodiment of any of the foregoing embodiments, the geometrically segmented coating section is on the leading end of the seal arc-segment and the first circumferential side of the seal arc-segment, and the second circumferential side excludes the geometrically segmented coating.
A gas turbine engine according to an example of the present disclosure includes a compressor section, a combustor in fluid communication with the compressor section, and a turbine section in fluid communication with the combustor. The turbine section has a blade outer air seal that has a seal arc-segment defining a gaspath side, a non-gaspath side, leading and trailing ends, and first and second circumferential sides. A portion of the gaspath side has a geometrically segmented coating section. The geometrically segmented coating section has a wall that has an array of cells, and a coating is disposed in the array of cells.
In a further embodiment of any of the foregoing embodiments, the geometrically segmented coating section is on at least one of the leading end of the seal arc-segment or the first circumferential side of the seal arc-segment.
In a further embodiment of any of the foregoing embodiments, a remaining portion of the gaspath side includes a non-segmented coating.
In a further embodiment of any of the foregoing embodiments, the array of cells includes elongated grooves.
In a further embodiment of any of the foregoing embodiments, the array of cells includes sloped grooves.
In a further embodiment of any of the foregoing embodiments, the array of cells includes at least one of parallel elongated grooves, sloped grooves, intersecting grooves, cylindrical cells, or polygonal cells.
The various features and advantages of the present disclosure will become apparent to those skilled in the art from the following detailed description. The drawings that accompany the detailed description can be briefly described as follows.
The exemplary 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 first (or low) pressure compressor 44 and a first (or 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 a 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 second (or high) pressure compressor 52 and a second (or high) pressure turbine 54. A combustor 56 is arranged in exemplary gas turbine 20 between the high pressure compressor 52 and the high pressure turbine 54. A mid-turbine frame 57 of the engine static structure 36 may be arranged generally between the high pressure turbine 54 and the low pressure turbine 46. The mid-turbine frame 57 further supports 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 then the high pressure compressor 52, mixed and burned with fuel in the combustor 56, then expanded over the high pressure turbine 54 and low pressure turbine 46. The mid-turbine frame 57 includes airfoils 59 which are in the core airflow path C. The turbines 46, 54 rotationally drive the respective low speed spool 30 and high speed spool 32 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, gear system 48 may be located aft of the low pressure compressor, or aft of the combustor section 26 or even aft of turbine section 28, and fan 42 may be positioned forward or aft of the location of gear system 48.
The engine 20 in one example is a high-bypass geared aircraft engine. In a further example, the 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 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 and less than about 5:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention 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 engine 20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and 35,000 ft (10,668 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 meters/second).
Blade outer air seals in general may include barrier coatings (e.g., thermal or environmental barrier coatings) on the gaspath side that serve to protect the underlying structure, typically formed of an alloy. A problem can be that the barrier coating spalls, leaving the underlying alloy exposed. In this regard, a portion of the gaspath side 66a of the disclosed seal arc-segment 66 has a geometrically segmented coating section 70 (“GSC section 70”). In particular, the seal arc-segment 66 includes the GSC section 70 in only one or more selected regions which are especially susceptible to spallation.
Two regions on blade outer air seals that may be more susceptible to spallation than other regions are the leading end and the circumferential side that the blade 64, when rotating, first encounters, which may be known as the blade arrival side. In this regard, the seal arc-segment 66 may include GSC sections 70 on the leading end 66c, on the first circumferential side 66e (the blade arrival side), or both. In the illustrated example, the seal arc-segment 66 has GSC sections 70 on both the leading end 66a and the first circumferential side 66e. The GSC sections 70 provide enhanced protection in those areas, thus facilitating improved spallation durability of the seal arc-segment 66.
The remainder of the gas path side 66a has a barrier coating 72 such as a ceramic coating. A ceramic material is a compound of metallic or metalloid elements bonded with nonmetallic elements or metalloid elements primarily in ionic or covalent bonds. Example ceramic materials may include, but are not limited to, oxides, carbides, nitrides, borides, silicides, and combinations thereof. Additional, example barrier coatings 72 may include, but are not limited to, ceramic coating systems, such as yttria stabilized with zirconia, hafnia, and/or gadolinia, gadolinia zirconate, molybdate, alumina, or combinations thereof, with or without underlying MCrAlY bond coats, where the M includes at least one of nickel, cobalt, iron, or combinations thereof.
The wall 74 may be formed of an alloy. Example alloys may include, but are not limited to, nickel alloys, cobalt alloys, a nickel alloy coated with cobalt or cobalt alloy, or non-nickel alloys that do not substantially react with ceramic.
The cells 76 facilitate reducing internal stresses in the coating 78 that may occur from sintering at relatively high surface temperatures during use in the engine 20. The sintering may result in partial melting, densification, and diffusional shrinkage of the coating 78 and thereby induce internal stresses. The cells 76 serve to produce faults in the coating 78. A fault is a crack or localized line or region in the coating 78 that is weaker than the surrounding regions. As an example, the faults may be regions where the coating 78 has localized lower density, i.e., higher porosity, than the surrounding regions, resulting in an inherent weakness at which the coating 78 can preferentially crack to release energy associated with the internal stresses (e.g., reducing shear and radial stresses). That is, the energy associated with the internal stresses may be dissipated in the faults such that there is less energy available for causing delamination cracking between the coating 78 and the underlying wall 74. As an example, the faults may emanate from the relatively sharp corners of the cells 76.
The GSC section 70 may be formed using several different fabrication techniques. As an example, the wall 74 may be fabricated by investment casting, additive manufacturing, brazing, or combinations thereof, but is not limited to such techniques. For instance, the cells 76 can be separately fabricated and brazed to the remaining portion of the wall 74, which can be investment cast or additively fabricated. Alternatively, the cells 76 can be formed by other techniques, such as depositing an alloy coating and removing sections of the alloy coating by machining, electro-discharge machining (EDM), or other removal process.
To produce the coating 78, ceramic coating material is deposited over the cells 76. The deposition process can include, but is not limited to, plasma spray or physical vapor deposition. There may be voids or pores in the coating 78; however, the coating 78 may substantially fully dense. For instance, the coating 78 may have a porosity of less than 15%.
The ceramic coating material fills or substantially fills the cells 76 and is deposited in a thickness that is greater than the height of the cells 76. At this stage, the surface of the coating 78 may have contours from the underlying cells 76. If such contours are undesired, the surface may be machined, ground, or abraded flat. For instance, the surface is reduced down to or close to the tops of the cells 76.
The cells 76 may be provided in a variety of different patterns. The figures that follow illustrate examples of such patterns. The examples are numbered with like reference numerals to designate like elements where appropriate, and reference numerals with the addition of one-hundred or multiples thereof designate modified elements that are understood to incorporate the same features and benefits of the corresponding elements.
Although a combination of features is shown in the illustrated examples, not all of them need to be combined to realize the benefits of various embodiments of this disclosure. In other words, a system designed according to an embodiment of this disclosure will not necessarily include all of the features shown in any one of the Figures or all of the portions schematically shown in the Figures. Moreover, selected features of one example embodiment may be combined with selected features of other example embodiments.
The preceding description is exemplary rather than limiting in nature. Variations and modifications to the disclosed examples may become apparent to those skilled in the art that do not necessarily depart from this disclosure. The scope of legal protection given to this disclosure can only be determined by studying the following claims.
