Blade outer air seal with machinable coating

A gas turbine engine includes a compressor section, a combustor section and a turbine section for rotation on an axis. The turbine section includes at least one row of rotating turbine blades each having a radially outer tip. A blade outer air seal is positioned radially outwardly of the radially outer tip. The blade outer air seal has a central web positioned radially outwardly of the radially outer tip. The blade outer air seal has an upstream mount arm and a downstream mount arm receiving mount structure from a static structure. The static structure has sealing members engaging an upstream outer surface of the upstream mount arm at an upstream seal material and a downstream outer surface of the downstream mount arm at a downstream seal material. A method is also disclosed.

BACKGROUND OF THE INVENTION

This application relates to the use of machinable coatings to provide sealing surfaces on a blade outer air seal.

Gas turbine engines typically include a fan delivering air into a bypass duct as propulsion air, and into a core engine. The core engine air moves into a compressor section where it is compressed and delivered into a combustor. The air is mixed with fuel and ignited in the combustor and passed downstream over turbine rotors driving them to rotate. The turbine rotors in turn rotate the fan and compressor rotors.

Improving the efficiency of gas turbine engines is important. To maximize the volume of the products of combustion passing over the turbine rotors, a blade outer air seal (“BOAS”) is placed radially outwardly of turbine blades to block the flow of products of combustion from avoiding the turbine blades.

There is a need for BOAS with improved temperature, environmental, and wear resistance for use in the operating conditions in a gas turbine engine.

SUMMARY OF THE INVENTION

A blade outer air seal according to an exemplary embodiment of this disclosure, among other possible things includes a center web having a first surface and a second surface opposite the first surface, and first and second faces joining the first surface to the second surface. The blade outer air seal also includes first and second mounting arms extending from the second surface. The second surface has a center portion between the first and second mounting arms and first and second side portions on either side of the center portion. The blade outer air seal also includes a continuous abradable coating disposed on the first surface of the center web, at least one of the first and second faces, and at least one of the first and second side portions of the second surface.

In a further example of the foregoing, the continuous abradable coating is a first continuous coating segment. Each of the mounting arms have an inner surface facing the center portion of the second surface and an outer surface opposite from the inner surface. The blade outer air seal also includes a second continuous abradable coating segment disposed on the center portion of the second surface and the inner surfaces of each of the mounting arms.

In a further example of any of the foregoing, the blade outer air seal comprises a ceramic matrix composite material.

In a further example of any of the foregoing, the blade outer air seal comprises a monolithic ceramic material.

In a further example of any of the foregoing, the coating comprises at least one of rare earth silicates, alkaline earth silicates, alkaline earth aluminosilicates, yttria-stabilized zirconia, alumina-stabilized zirconia, mullite, titania, chromia, silicon, silicon oxides, silicon carbides, silicon oxycarbides, barium-magnesium aluminosilicate, hafnium oxides such as hafnon, hafnium silicon oxides, alumina-stabilized zirconia, zirconium oxides such as zircon, yttrium oxides such as yttria, mullite, and combinations thereof.

A gas turbine engine according to an exemplary embodiment of this disclosure, among other possible things includes a turbine section arranged along a central engine axis. The turbine section has a turbine with at least one blade rotatable around the central engine axis. The at least one blade has a tip. The gas turbine engine also includes at least one blade outer air seal arranged radially outward from the tip and attached to an engine static structure. The blade outer air seal includes a web having a radially inner surface adjacent the tip and a radially outer surface, a forward face and an axial face joining the radially inner surface and radially outer surface. First and second mounting arms extend radially outward from the radially outer surface. The radially outer surface of the web has a forward portion, an aft portion, and a center portion between the first and second mounting arms. The blade outer air seal also includes a continuous abradable coating disposed on the radially inner surface of the web, and at least one of the forward portion and aft portion of the radially outer surface.

In a further example of the foregoing, the continuous abradable coating is disposed on the forward portion of the radially outer surface and the forward face of the web.

In a further example of any of the foregoing, the continuous abradable coating is a first continuous coating segment. The gas turbine engine also includes a second continuous coating segment disposed on an aft surface of the first mounting arm, a forward surface of the second mounting arm, and the center portion of the radially outer surface.

In a further example of any of the foregoing, the blade outer air seal comprises a ceramic matrix composite material.

In a further example of any of the foregoing, the continuous abradable coating is disposed on the forward portion of the radially outer surface and the forward face of the web.

In a further example of any of the foregoing, the blade outer air seal comprises a ceramic matrix composite material.

In a further example of any of the foregoing, the coating comprises at least one of rare earth silicates, alkaline earth silicates, alkaline earth aluminosilicates, yttria-stabilized zirconia, alumina-stabilized zirconia, mullite, titania, chromia, silicon, silicon oxides, silicon carbides, silicon oxycarbides, barium-magnesium aluminosilicate, hafnium oxides such as hafnon, hafnium silicon oxides, alumina-stabilized zirconia, zirconium oxides such as zircon, yttrium oxides such as yttria, mullite, and combinations thereof.

In a further example of any of the foregoing, the coating comprises at least one of rare earth silicates, alkaline earth silicates, alkaline earth aluminosilicates, yttria-stabilized zirconia, alumina-stabilized zirconia, mullite, titania, chromia, silicon, silicon oxides, silicon carbides, silicon oxycarbides, barium-magnesium aluminosilicate, hafnium oxides such as hafnon, hafnium silicon oxides, alumina-stabilized zirconia, zirconium oxides such as zircon, yttrium oxides such as yttria, mullite, and combinations thereof.

A method of coating a blade outer air seal according to an exemplary embodiment of this disclosure, among other possible things includes applying an abradable coating to a blade outer air seal. The blade outer air seal includes a center web having a first surface, a second surface opposite the first surface, and first and second faces joining the first surface to the second surface. The blade outer air seal also includes first and second mounting arms extending from the second surface. The second surface has a center portion between the first and second mounting arms and first and second side portions on either side of the center portion. The abradable coating forms a first continuous coating segment on the first surface of the center web, at least one of the first and second faces, and at least one of the first and second side portions of the second surface.

In a further example of the foregoing, each of the mounting arms have an inner surface facing the center portion of the second surface and an outer surface opposite from the inner surface. The blade outer air seal also includes a second continuous abradable coating segment disposed on the center portion of the second surface and the inner surfaces of each of the mounting arms.

In a further example of any of the foregoing, the second coating segment is applied simultaneously with the first coating segment.

In a further example of any of the foregoing, the blade outer air seal comprises a ceramic matrix composite material.

In a further example of any of the foregoing, the blade outer air seal comprises a monolithic ceramic material.

In a further example of any of the foregoing, the coating comprises at least one of rare earth silicates, alkaline earth silicates, alkaline earth aluminosilicates, yttria-stabilized zirconia, alumina-stabilized zirconia, mullite, titania, chromia, silicon, silicon oxides, silicon carbides, silicon oxycarbides, barium-magnesium aluminosilicate, hafnium oxides such as hafnon, hafnium silicon oxides, alumina-stabilized zirconia, zirconium oxides such as zircon, yttrium oxides such as yttria, mullite, and combinations thereof.

In a further example of any of the foregoing, the coating segment is applied by plasma spraying, slurry infiltration and sintering, or melt infiltration.

DETAILED DESCRIPTION

FIG.1schematically illustrates a gas turbine engine20. The gas turbine engine20is disclosed herein as a two-spool turbofan that generally incorporates a fan section22, a compressor section24, a combustor section26and a turbine section28. The fan section22may include a single-stage fan42having a plurality of fan blades43. The fan blades43may have a fixed stagger angle or may have a variable pitch to direct incoming airflow from an engine inlet. The fan42drives air along a bypass flow path B in a bypass duct13defined within a housing15such as a fan case or nacelle, and also drives air along a core flow path C for compression and communication into the combustor section26then expansion through the turbine section28. A splitter29aft of the fan42divides the air between the bypass flow path B and the core flow path C. The housing15may surround the fan42to establish an outer diameter of the bypass duct13. The splitter29may establish an inner diameter of the bypass duct13. Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with two-spool turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures. The engine20may incorporate a variable area nozzle for varying an exit area of the bypass flow path B and/or a thrust reverser for generating reverse thrust.

The low speed spool30generally includes an inner shaft40that interconnects, a first (or low) pressure compressor44and a first (or low) pressure turbine46. The inner shaft40is connected to the fan42through a speed change mechanism, which in the exemplary gas turbine engine20is illustrated as a geared architecture48to drive the fan42at a lower speed than the low speed spool30. The inner shaft40may interconnect the low pressure compressor44and low pressure turbine46such that the low pressure compressor44and low pressure turbine46are rotatable at a common speed and in a common direction. In other embodiments, the low pressure turbine46drives both the fan42and low pressure compressor44through the geared architecture48such that the fan42and low pressure compressor44are rotatable at a common speed. Although this application discloses geared architecture48, its teaching may benefit direct drive engines having no geared architecture. The high speed spool32includes an outer shaft50that interconnects a second (or high) pressure compressor52and a second (or high) pressure turbine54. A combustor56is arranged in the exemplary gas turbine20between the high pressure compressor52and the high pressure turbine54. A mid-turbine frame57of the engine static structure36may be arranged generally between the high pressure turbine54and the low pressure turbine46. The mid-turbine frame57further supports bearing systems38in the turbine section28. The inner shaft40and the outer shaft50are concentric and rotate via bearing systems38about the engine central longitudinal axis A which is collinear with their longitudinal axes.

The fan42may have at least 10 fan blades43but no more than 20 or 24 fan blades43. In examples, the fan42may have between 12 and 18 fan blades43, such as 14 fan blades43. An exemplary fan size measurement is a maximum radius between the tips of the fan blades43and the engine central longitudinal axis A. The maximum radius of the fan blades43can be at least 40 inches, or more narrowly no more than 75 inches. For example, the maximum radius of the fan blades43can be between 45 inches and 60 inches, such as between 50 inches and 55 inches. Another exemplary fan size measurement is a hub radius, which is defined as distance between a hub of the fan42at a location of the leading edges of the fan blades43and the engine central longitudinal axis A. The fan blades43may establish a fan hub-to-tip ratio, which is defined as a ratio of the hub radius divided by the maximum radius of the fan42. The fan hub-to-tip ratio can be less than or equal to 0.35, or more narrowly greater than or equal to 0.20, such as between 0.25 and 0.30. The combination of fan blade counts and fan hub-to-tip ratios disclosed herein can provide the engine20with a relatively compact fan arrangement.

The low pressure compressor44, high pressure compressor52, high pressure turbine54and low pressure turbine46each include one or more stages having a row of rotatable airfoils. Each stage may include a row of vanes adjacent the rotatable airfoils. The rotatable airfoils are schematically indicated at47, and the vanes are schematically indicated at49.

The low pressure compressor44and low pressure turbine46can include an equal number of stages. For example, the engine20can include a three-stage low pressure compressor44, an eight-stage high pressure compressor52, a two-stage high pressure turbine54, and a three-stage low pressure turbine46to provide a total of sixteen stages. In other examples, the low pressure compressor44includes a different (e.g., greater) number of stages than the low pressure turbine46. For example, the engine20can include a five-stage low pressure compressor44, a nine-stage high pressure compressor52, a two-stage high pressure turbine54, and a four-stage low pressure turbine46to provide a total of twenty stages. In other embodiments, the engine20includes a four-stage low pressure compressor44, a nine-stage high pressure compressor52, a two-stage high pressure turbine54, and a three-stage low pressure turbine46to provide a total of eighteen stages. It should be understood that the engine20can incorporate other compressor and turbine stage counts, including any combination of stages disclosed herein.

The engine20may be a high-bypass geared aircraft engine. The bypass ratio can be greater than or equal to 10.0 and less than or equal to about 18.0, or more narrowly can be less than or equal to 16.0. The geared architecture48may be an epicyclic gear train, such as a planetary gear system or a star gear system. The epicyclic gear train may include a sun gear, a ring gear, a plurality of intermediate gears meshing with the sun gear and ring gear, and a carrier that supports the intermediate gears. The sun gear may provide an input to the gear train. The ring gear (e.g., star gear system) or carrier (e.g., planetary gear system) may provide an output of the gear train to drive the fan42. A gear reduction ratio may be greater than or equal to 2.3, or more narrowly greater than or equal to 3.0, and in some embodiments the gear reduction ratio is greater than or equal to 3.4. The gear reduction ratio may be less than or equal to 4.0. The fan diameter is significantly larger than that of the low pressure compressor44. The low pressure turbine46can have a pressure ratio that is greater than or equal to 8.0 and in some embodiments is greater than or equal to 10.0. The low pressure turbine pressure ratio can be less than or equal to 13.0, or more narrowly less than or equal to 12.0. Low pressure turbine46pressure ratio is pressure measured prior to an inlet of low pressure turbine46as related to the pressure at the outlet of the low pressure turbine46prior to an exhaust nozzle. 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. All of these parameters are measured at the cruise condition described below.

A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section22of the engine20is 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 1 bm of fuel being burned divided by 1 bf of thrust the engine produces at that minimum point. The engine parameters described above, and those in the next paragraph are measured at this condition unless otherwise specified.

“Fan pressure ratio” is the pressure ratio across the fan blade43alone, without a Fan Exit Guide Vane (“FEGV”) system. A distance is established in a radial direction between the inner and outer diameters of the bypass duct13at an axial position corresponding to a leading edge of the splitter29relative to the engine central longitudinal axis A. The fan pressure ratio is a spanwise average of the pressure ratios measured across the fan blade43alone over radial positions corresponding to the distance. The fan pressure ratio can be less than or equal to 1.45, or more narrowly greater than or equal to 1.25, such as between 1.30 and 1.40. “Corrected fan tip speed” is the actual fan tip speed in ft/see divided by an industry standard temperature correction of [(Tram ° R)/(518.7° R)]0.5. The corrected fan tip speed can be less than or equal to 1150.0 ft/second (350.5 meters/second), and can be greater than or equal to 1000.0 ft/second (304.8 meters/second).

The fan42, low pressure compressor44and high pressure compressor52can provide different amounts of compression of the incoming airflow that is delivered downstream to the turbine section28and cooperate to establish an overall pressure ratio (OPR). The OPR is a product of the fan pressure ratio across a root (i.e., 0% span) of the fan blade43alone, a pressure ratio across the low pressure compressor44and a pressure ratio across the high pressure compressor52. The pressure ratio of the low pressure compressor44is measured as the pressure at the exit of the low pressure compressor44divided by the pressure at the inlet of the low pressure compressor44. In examples, a sum of the pressure ratio of the low pressure compressor44and the fan pressure ratio is between 3.0 and 6.0, or more narrowly is between 4.0 and 5.5. The pressure ratio of the high pressure compressor ratio52is measured as the pressure at the exit of the high pressure compressor52divided by the pressure at the inlet of the high pressure compressor52. In examples, the pressure ratio of the high pressure compressor52is between 9.0 and 12.0, or more narrowly is between 10.0 and 11.5. The OPR can be equal to or greater than 45.0, and can be less than or equal to 70.0, such as between 50.0 and 60.0. The overall and compressor pressure ratios disclosed herein are measured at the cruise condition described above, and can be utilized in two-spool architectures such as the engine20as well as three-spool engine architectures.

The engine20establishes a turbine entry temperature (TET). The TET is defined as a maximum temperature of combustion products communicated to an inlet of the turbine section28at a maximum takeoff (MTO) condition. The inlet is established at the leading edges of the axially forwardmost row of airfoils of the turbine section28, and MTO is measured at maximum thrust of the engine20at static sea-level and 86 degrees Fahrenheit (° F.). The TET may be greater than or equal to 2700.0° F., or more narrowly less than or equal to 3500.0° F., such as between 2750.0° F. and 3350.0° F. The relatively high TET can be utilized in combination with the other techniques disclosed herein to provide a compact turbine arrangement.

The engine20establishes an exhaust gas temperature (EGT). The EGT is defined as a maximum temperature of combustion products in the core flow path C communicated to at the trailing edges of the axially aftmost row of airfoils of the turbine section28at the MTO condition. The EGT may be less than or equal to 1000.0° F., or more narrowly greater than or equal to 800.0° F., such as between 900.0° F. and 975.0° F. The relatively low EGT can be utilized in combination with the other techniques disclosed herein to reduce fuel consumption.

FIG.2shows a turbine section28having rotating turbine blades102with a radially outer tip103. A vane is positioned upstream of the turbine blade102(not shown inFIG.2). In one example the turbine blade102is a blade of the high pressure turbine54(FIG.1). A blade outer air seal (BOAS)105is positioned radially outwardly of the tip103with respect to a central engine axis. A plurality of BOAS (not shown) are arranged circumferentially round the central engine axis.

The BOAS105is formed out of ceramic matrix composite materials (“CMCs”) or a monolithic ceramic. A CMC material is comprised of one or more ceramic reinforcement plies in a ceramic matrix. Example ceramic matrices are silicon-containing ceramic, such as but not limited to, a silicon carbide (SiC) matrix or a silicon nitride (Si3N4) matrix. Example ceramic reinforcement of the CMC are silicon-containing ceramic fibers, such as but not limited to, silicon carbide (SiC) fiber or silicon nitride (Si3N4) fibers. The CMC may be, but is not limited to, a SiC/SiC ceramic matrix composite in which SiC fiber plies are disposed within a SiC matrix. A fiber ply has a fiber architecture, which refers to an ordered arrangement of the fiber tows relative to one another, such as a 2D woven ply or a 3D structure. A monolithic ceramic does not contain fibers or reinforcements and is formed of a single material. Example monolithic ceramics include silicon-containing ceramics, such as silicon carbide (SiC) or silicon nitride (Si3N4).

The static structure106may be an engine20casing structure.

The BOAS105has two mount arms108a/108bwhich extend generally perpendicular from a web110. A radially inner (with respect to the central engine axis A) surface111of the center web110is arranged adjacent the tip103of the turbine blade102when the BOAS105is attached to the static structure106. As such, the radially inner surface111is in the gaspath of the engine20. The radially inner surface111may include one or more coatings to enhance its wear, temperature, and environmental resistance. The coatings can include environmental barrier coatings or thermal barrier coatings, examples of which are known in the art. In addition, the radially inner surface111may include an abradable coating, which can accommodate rubbing of the tip103against the BOAS105.

The BOAS105has other, non-gaspath surfaces which are not exposed to the engine gaspath. For instance, the web110has a radially outer surface112, which includes a forward portion112a, an aft portion112b, and a center portion112cbetween the mounting arms108a/108b. In addition, the mounting arms108a/108b, which extend in a radially outer direction from the radially outer surface112, e.g., away from the gaspath, each have forward and aft surfaces109a/109b, respectively. The aft surface109bof the mounting arm108aand the forward surface109aof the mounting arm108bface inward towards the center portion112cwhile the forward surface109aof the mounting arm108aand the aft surface109bof the mounting arm108bface outward with respect to the center portion112c. In conventional BOAS105, these non-gaspath surfaces do not include coatings.

However, it has been discovered that application of at least an abradable or machinable coating to one or more of these non-gaspath surfaces can improve the lifetime of the BOAS105as will be discussed herein. Other coatings such as thermal barrier coatings or environmental barrier coatings, which are known in the art, could be used together with the abradable coating. Though the non-gaspath surfaces are not subject to abrasion as the radially inner surface111is due to possible tip103rub, the abradable coating is relatively easy to apply, cost-effective, and provides a measure of thermal protection to the BOAS105. Specifically, the BOAS105is subject to a temperature gradient because the radially inner surface111is subject to hot gaspath temperatures while the radially outer surface112and mounting arms108a/108bare not.

To that end,FIG.2Bshows a coated BOAS105. The BOAS105includes an abradable coating200on the radially inner surface111, at least a portion of the radially outer surface112, and optionally, the forward and aft surfaces109a/109bof the mounting arms108a/108b. The coating200may be or include, for example, rare earth silicates, alkaline earth silicates, alkaline earth aluminosilicates, yttria-stabilized zirconia, alumina-stabilized zirconia, mullite, titania, chromia, silicon, silicon oxides, silicon carbides, silicon oxycarbides, barium-magnesium aluminosilicate, hafnium oxides such as hafnon, hafnium silicon oxides, alumina-stabilized zirconia, zirconium oxides such as zircon, yttrium oxides such as yttria, mullite, and combinations thereof. In a particular example, the abradable coating200includes at least one of hafnon, zircon, and mullite.

Application of the abradable coating200on the non-gaspath surfaces reduces the temperature gradient across the BOAS105. For at least a CMC BOAS105, this in turn reduces interlaminar stresses within the BOAS and improves its longevity.

The coating200is applied in any suitable way known in the art, such as air plasma spraying, slurry infiltration and sintering, or melt infiltration.

In one example, the coating200is continuous along the radially inner surface111and at least a portion of the radially outer surface112. That is, the coating200wraps around at least one axial face of the BOAS, such as the forward axial face113a, the aft axial face113b, or both and extends along the radially inner surface111and at least a portion of the radially outer surface.

In the particular example shown inFIG.2, the coating200includes two coating sections200a/200b. The first coating section200ais a continuous coating that coats the radially inner surface111, forward axial face113a, and forward portion112aof the radially outer face of the center web110. The second coating section200bis a continuous coating that coats the center portion112cof the radially outer face, aft surface109bof mounting arm108a, and forward surface of109aof mounting arm108b. In some examples, the coating sections200aand200bare applied to the BOAS simultaneously according to any of the example methods discussed above.

As used herein, the term “about” and “approximately” have the typical meanings in the art, however in a particular example “about” and “approximately” can mean deviations of up to 10% of the values described herein.

Although the different examples are illustrated as having specific components, the examples of this disclosure are not limited to those particular combinations. It is possible to use some of the components or features from any of the embodiments in combination with features or components from any of the other embodiments.