Patent Description:
Gas turbine engines, and other turbomachines, include multiple sections, such as a fan section, a compressor section, a combustor section, a turbine section, and an exhaust section. Air moves into the engine through the fan section. Airfoil arrays in the compressor section rotate to compress the air, which is then mixed with fuel and combusted in the combustor section. The products of combustion are expanded to rotatably drive airfoil arrays in the turbine section. Rotating the airfoil arrays in the turbine section drives rotation of the fan and compressor sections.

A blade outer air seal (BOAS) array includes blade outer air seal (BOAS) segments circumferentially disposed about at least a portion of the airfoil arrays. As known, the blade outer air seal environment is exposed to temperature extremes and other harsh environmental conditions, which may affect the integrity of the blade outer air seal segments. In addition, high relative movements/displacements between the BOAS segment/array (an exemplary gas turbine engine component) and surrounding static hardware (e.g., stator vanes) due to the varying thermal environment in the operational temperature range may, in particular, expose a leading edge portion of the BOAS to high heat loads, potentially shortening BOAS life and/or compelling additional cooling flow. The leading edge portion of other gas turbine engine components may also be exposed to high heat loads due to high relative movements/displacements between the static hardware and the gas turbine engine component, potentially shortening the life of the gas turbine engine component and/or compelling additional cooling flow.

A prior art gas turbine engine component having the features of the preamble to claim <NUM> is disclosed in <CIT>. Another prior art gas turbine component having a curved transition region is disclosed in <CIT>, which forms prior art under Article <NUM>(<NUM>) EPC.

A prior art BOAS with a leading edge wall and cooling passage having an undercut profile is disclosed in <CIT>.

The present invention provides a gas turbine engine component in accordance with claim <NUM>.

In any of the foregoing embodiments, a static structure is configured to be disposed adjacent and upstream of the gas turbine engine component in a gas turbine engine and each of the static structure and the gas turbine engine component is configured to move relative to each other because of thermal or mechanical deflections. The elongated transition portion has an axial length that is greater than a radial height by up to one order of magnitude. The axial length of the elongated transition portion is about three to about ten times the radial height of the elongated transition portion. The elongated transition portion has a chamfer of less than about <NUM> degrees combined with a radius. The elongated transition portion is configured as a chamfer blended with a radius to at least one of the leading edge or the proximate flowpath surface of the main body. The gas turbine engine component comprises a blade outer air seal (BOAS) segment.

The present disclosure will become more fully understood from the detailed description and the accompanying drawings wherein:.

The detailed description of exemplary embodiments herein makes reference to the accompanying drawings, which show exemplary embodiments by way of illustration and its best mode, and not of limitation. While these exemplary embodiments are described in sufficient detail to enable those skilled in the art to practice the invention, it should be understood that other embodiments may be realized and that logical, chemical and mechanical changes may be made without departing from the scope of the invention according to the claims. 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. Moreover, many of the functions or steps may be outsourced to or performed by one or more third parties. Also, any reference to attached, fixed, connected or the like may include permanent, removable, temporary, partial, full and/or any other possible attachment option. It is to be understood that unless specifically stated otherwise, references to "a," "an," and/or "the" may include one or more than one and that reference to an item in the singular may also include the item in the plural.

Various embodiments are directed to gas turbine engines and gas turbine engine components such as blade outer air seal (BOAS) segments with optimized leading edge geometry. Relative movement or shifts due to the varying thermal environment between a non-rotating component and an adjacent BOAS in a turbine or compressor stage of a gas turbine engine can result in a leading edge portion of the BOAS projecting into the hot core flowpath of the gas turbine engine, resulting in a high heat load for the BOAS leading edge portion, thereby shortening BOAS life and/or compelling additional cooling. Various embodiments permit the hot core flowpath air to impinge on the BOAS leading edge portion at a reduced incidence angle (relative to conventional leading edge geometry), thereby minimizing exposure of the BOAS leading edge portion to high heat transfer coefficients from the hot core flowpath air and thus extending BOAS life and/or minimizing cooling requirements. While a BOAS segment having a leading edge portion with an optimized geometry is described herein, it is to be understood that the BOAS segment is an exemplary gas turbine engine component and that other gas turbine engine components may benefit from an optimized leading edge geometry according to various embodiments.

According to various embodiments, and with reference to <FIG>, a gas turbine engine <NUM> is schematically illustrated. According to various embodiments, gas turbine engine <NUM> may be a two-spool turbofan that generally incorporates a fan section <NUM>, a compressor section <NUM>, a combustor section <NUM> and a turbine section <NUM>, for example. Alternative engines might include an augmentor section (not shown) among other systems or features, according to various embodiments. According to various embodiments, the fan section <NUM> drives air along a bypass flowpath B while the compressor section <NUM> drives air along a core flowpath for compression and communication into the combustor section <NUM> then expansion through the turbine section <NUM>. Although depicted as a 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 turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures, non-geared turbine engines, and land-based turbines, according to various embodiments.

According to various embodiments, gas turbine engine <NUM> may generally include a first spool <NUM> and a second spool <NUM> mounted for rotation about an engine central axis A relative to an engine static structure <NUM> via several bearing systems <NUM>. It should be understood that various bearing systems <NUM> at various locations may alternatively or additionally be provided, according to various embodiments. According to various embodiments, the first spool <NUM> may generally include a first shaft <NUM> that interconnects a fan <NUM>, a first compressor <NUM> and a first turbine <NUM>. According to various embodiments, the first shaft <NUM> may be connected to the fan <NUM> through a gear assembly of a fan drive gear system <NUM> to drive the fan <NUM> at a lower speed than the first spool <NUM>. According to various embodiments, the second spool <NUM> may include a second shaft <NUM> that interconnects a second compressor <NUM> and second turbine <NUM>. According to various embodiments, the first spool <NUM> may run at a relatively lower pressure than the second spool <NUM>. It is to be understood that "low pressure" and "high pressure" or variations thereof as used herein are relative terms indicating that the high pressure is greater than the low pressure. According to various embodiments, an annular combustor <NUM> may be arranged between the second compressor <NUM> and the second turbine <NUM>. According to various embodiments, the first shaft <NUM> and the second shaft <NUM> may be concentric and rotate via bearing systems <NUM> about the engine central axis A which is collinear with their longitudinal axes, according to various embodiments.

According to various embodiments, the core airflow may be compressed by the first compressor <NUM> then the second compressor <NUM>, mixed and burned with fuel in the annular combustor <NUM>, then expanded over the second turbine <NUM> and first turbine <NUM>. According to various embodiments, the first turbine <NUM> and the second turbine <NUM> may rotationally drive, respectively, the first spool <NUM> and the second spool <NUM> in response to the expansion. According to various embodiments, gas turbine engine <NUM> may be a high-bypass geared aircraft engine that has a bypass ratio that is greater than about six (<NUM>), with an example embodiment being greater than ten (<NUM>). According to various embodiments, the gear assembly of the fan drive gear system <NUM> may be an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about <NUM>:<NUM> and the first turbine <NUM> may have a pressure ratio that is greater than about <NUM>, for example. According to various embodiments, the first turbine <NUM> pressure ratio is pressure measured prior to inlet of first turbine <NUM> as related to the pressure at the outlet of the first turbine <NUM> prior to an exhaust nozzle. According to various embodiments, first turbine <NUM> may have a maximum rotor diameter and the fan <NUM> may have a fan diameter such that a ratio of the maximum rotor diameter divided by the fan diameter is less than <NUM>. It should be understood, however, that the above parameters are only exemplary.

A significant amount of thrust may be provided by the bypass flow B due to the high bypass ratio. According to various embodiments, the fan section <NUM> of the gas turbine engine <NUM> may be designed for a particular flight condition-typically cruise at an airspeed of <NUM> Mach and altitude of <NUM>,<NUM> feet (<NUM>). The flight condition of <NUM> Mach and <NUM>,<NUM> feet (<NUM>) may be a condition at which an engine is operating at its best fuel consumption. To make an accurate comparison of fuel consumption between engines, fuel consumption is reduced to a common metric which is applicable to all types and sizes of turbojets and turbofans. The term that may be used to compare fuel consumption between engines is thrust specific fuel consumption, or TSFC. This is an engine's fuel consumption in pounds per hour divided by the net thrust. Stated another way, TSFC is the amount of fuel required to produce one pound of thrust. The TSFC unit is pounds per hour per pounds of thrust (lb/hr/lb Fn). When the reference is to a turbojet or turbofan engine, TSFC is often simply called specific fuel consumption, or SFC. "Low fan pressure ratio" is the pressure ratio across the fan blade alone, without a fan exit guide vane system. "Low corrected fan tip speed" is the actual fan tip speed in feet per second divided by an industry standard temperature correction of <MAT>. The "Low corrected fan tip speed" as disclosed herein according to one non-limiting embodiment may be less than about <NUM> feet per second (<NUM>/s).

Each of the first and second compressors <NUM> and <NUM> and first and second turbines <NUM> and <NUM> in the gas turbine engine <NUM> comprises interspersed stages of rotor blades <NUM> and stator vanes <NUM>. The rotor blades <NUM> rotate about the centerline with the associated shaft while the stator vanes <NUM> remain stationary about the centerline. The first and second compressors <NUM> and <NUM> in the gas turbine engine may each comprise one or more compressor stages. The first and second turbines <NUM> and <NUM> in the gas turbine engine <NUM> may each comprise one or more turbine stages. Each compressor stage and turbine stage may comprise multiple sets of rotating blades ("rotor blades") and stationary vanes ("stator vanes"). For example, <FIG> schematically shows, by example, a first turbine stage of the second turbine <NUM> (a high-pressure turbine (HPT)) in the turbine section of the gas turbine engine. Unless otherwise indicated, the term "blade stage" refers to at least one of a turbine stage or a compressor stage.

With continued reference to <FIG> and <FIG>, according to various embodiments, the depicted first turbine stage of the HPT comprises the rotor blade <NUM> and the stator vane <NUM>. The stator vane <NUM> may have an inner stator vane platform <NUM> and an outer stator vane platform <NUM>. The rotor blade <NUM> may comprise a blade airfoil section <NUM> and a platform <NUM>, such as a rotor blade platform. A blade outer air seal (BOAS) segment <NUM> is attached to an engine case structure <NUM> of the gas turbine engine <NUM> by a receiving portion <NUM> of the engine case structure <NUM>. The BOAS segment <NUM> faces the rotor blade <NUM> (an exemplary turbine blade in <FIG>) to define a radial tip clearance <NUM> between the rotor blade <NUM> and the BOAS segment <NUM> (more particularly, between a proximate inner diameter flowpath surface <NUM> of the BOAS and the turbine blade tip). A minimal radial tip clearance <NUM> is sought to be maintained as the smaller the clearance, the greater the turbine efficiency. The BOAS segment <NUM> locally bounds the radially outboard extreme of the core flowpath through the gas turbine engine <NUM>. Although only one BOAS segment <NUM> is shown in <FIG>, the turbine stage comprises an associated array of blade outer air seal segments. A number of BOAS segments <NUM> may be arranged circumferentially about engine axis A to form a shroud, according to various embodiments. According to various embodiments, the BOAS segments <NUM> may alternatively be formed as a unitary BOAS structure, with the same features described herein.

Still referring to <FIG> and now specifically to <FIG>, according to various embodiments, the BOAS segment <NUM> may include a main body <NUM> that extends generally axially from a leading edge portion <NUM> to a trailing edge portion <NUM> and from a radially outward facing surface <NUM> at an outboard side of BOAS segment <NUM> to the inner diameter flowpath surface <NUM> at an inboard side of BOAS segment <NUM>. The leading edge portion <NUM> of BOAS segment <NUM> includes a leading edge <NUM> and a leading edge wall 56a. In accordance with various embodiments, the leading edge wall 56a includes an elongated transition portion 56b that extends between the leading edge <NUM> and the inner diameter flowpath surface <NUM> of the BOAS segment as hereinafter described. The BOAS segment <NUM> also includes at least one leading attachment portion 60a (also referred to as "attachment portions 60a") disposed at or near the leading edge portion <NUM> and at least one trailing attachment portion 60b (also referred to as "attachment portions 60b") disposed at or near the trailing edge portion <NUM>. Each of the attachment portions 60a, 60b may define a flange <NUM>. Flange <NUM> of attachment portions 60a and/or 60b may extend in an axially aft direction. Flange <NUM> of attachment portions 60a and/or 60b may alternatively extend in an axially forward direction, as shown in the figures. Flange <NUM> of attachment portions 60a and/or 60b may alternatively extend in and/or out of the page. Each axially extending flange <NUM> corresponds to the receiving portion <NUM> of the engine case structure <NUM> to support and attach the BOAS segment <NUM> (shown schematically in <FIG>). According to various embodiments, the attachment portions 60a may be circumferentially offset, circumferentially aligned, or a combination of both, from the attachment portions 60b in response to BOAS segment <NUM> parameters.

Still referring to <FIG> and <FIG> and now to <FIG>, according to various embodiments, and as noted previously, the leading edge portion <NUM> of the BOAS segment has an elongated transition portion 56b extending from the leading edge <NUM> to the inner diameter flowpath surface <NUM>. The elongated transition portion 56b, which may have various configurations as hereinafter described, defines an optimized leading edge geometry that deviates away from a conventional leading edge geometry (see, e.g., <FIG>). The elongated transition portion 56b has an axial length (L) that is greater than a radial height (H), of up to one order of magnitude. The axial length of the elongated transition portion 56b is about three to about ten times the radial height of the elongated transition portion 56b. The axial length is defined as the length between a first tangency point <NUM> defined between the leading edge (face) <NUM> and the leading edge wall 56a and a second tangency point <NUM> defined between the leading edge wall 56a and the inner diameter flowpath surface <NUM>. In various embodiments, the radial height is defined as the radial distance between the inner diameter flowpath surface <NUM> and the first tangency point <NUM>.

Still referring to <FIG> and now specifically to <FIG>, according to the invention, the elongated transition portion 56b is configured as an ellipse. The ellipse may have an elliptical factor of greater than about <NUM>, wherein the elliptical factor is defined as a length of a major axis divided by the length of a minor axis. Referring now specifically to <FIG>, according to various embodiments, the elongated transition portion is configured with the ellipse (<FIG>) or, in arrangements outside the scope of the present invention, with a chamfer in combination with a radius (<FIG>). A degree of the chamfer is less than about <NUM> degrees. The shape of the elongated transition portion (the ellipse) in <FIG> is the same as the shape of the elongated transition portion in <FIG>.

As herein described, the leading edge portion (more particularly, the elongated transition portion) of the BOAS segment (and other gas turbine engine components) has a geometry such that over an operational temperature range, thermal and/or mechanical deflections of a non-rotating structure (e.g., the upstream stator vane <NUM> depicted in <FIG>) upstream of the BOAS array relative to thermal and/or mechanical deflections of the BOAS array cause relative movement of the non-rotating structure and the BOAS array to expose the leading edge portion of the BOAS to hot core flowpath air. The BOAS array may be radially deflected outboard of the upstream non-rotating structure as shown in <FIG> (a "waterfall condition") or radially deflected inboard of the upstream non-rotating structure as shown in <FIG> (a "dam condition").

More specifically, <FIG> depicts a BOAS segment/array with an untreated (i.e., not configured in accordance with various embodiments) leading edge portion. <FIG> depict a BOAS segment/array with a conventional transition between the leading edge wall and the inner diameter flowpath surface <NUM>, with <FIG> depicting a conventional chamfer and <FIG> depicting a conventional radius. As a result, the leading edge portion of the BOAS segment/array in each of <FIG> is fully exposed to hot core flowpath air ("hot gas flow"). By contrast, the BOAS segment/array in <FIG> has the elongated transition portion configured as the ellipse in accordance with the present invention, such that the hot core flowpath air is transitioned from the upstream vane to the BOAS array with a reduced incidence angle (relative to the conventional or untreated leading edge geometry) accommodating an increased range of relative radial deflection.

As noted previously, the BOAS array may alternatively be radially deflected inboard of the upstream non-rotating structure (a "dam condition) as shown in <FIG> depicts a BOAS segment/array with an untreated (i.e., not configured in accordance with various embodiments) leading edge portion. <FIG> depicts a BOAS segment/array with a conventional chamfered transition between the leading edge wall and the inner flowstream path. As a result of the relative radial deflection between the upstream non-rotating component (a stator vane in the depicted embodiment) and the adjacent BOAS segment/array, the leading edge portion of the BOAS segment/array in each of <FIG> is fully exposed to hot core flowpath air and high heat transfer coefficients. By contrast, the BOAS segment/array in <FIG> has the elongated transition portion configured as an ellipse in accordance with the present invention, so as to permit more deflection of the hot core flowpath air off the leading edge portion of the BOAS segment/array relative to an untreated or conventional BOAS segment/array (<FIG> and <FIG>). As noted previously, the BOAS segment/array in <FIG> has an elongated transition portion configured as a chamfer with a blended radius, so as to also permit more deflection of the hot core flowpath air off the leading edge portion of the BOAS segment/array relative to an untreated or conventional BOAS segment/array (<FIG> and <FIG>). The elongated transition portion configured as a chamfer with blended radii as depicted in <FIG> is also easier to make and inspect relative to other transitions. <FIG> is a schematic view of the leading edge portion of an exemplary BOAS segment outside the scope of the present invention illustrating that a transition of the elongated transition portion to the leading edge (face) of the BOAS segment may produce a corner rather than a tangent as in <FIG> and <FIG> is a schematic view of the leading edge portion of an exemplary BOAS segment outside the scope of the present invention illustrating an elongated transition portion comprising a short chamfer between two radii.

During gas turbine engine <NUM> operation, and over the operational temperature range, the BOAS segment <NUM> is subjected to different thermal loads and environmental conditions (i.e., the thermal environment surrounding each turbine or compressor stage varies during operation). As a result, the thermal and/or mechanical deflections of the non-rotating structure adjacent to the BOAS segment array and the thermal and/or mechanical deflections of the BOAS segment array may be such that relative movement exposes the leading edge portion to hot core flowpath air. According to various embodiments, the leading edge portion is configured such that the hot core flowpath air is transitioned from the upstream non-rotating structure (e.g., the upstream stator vane) to the BOAS array with a reduced incidence angle that accommodates an increased range of relative radial deflections. The variation in the radial clearance between the stationary vane and the adjacent BOAS is a result of how the outer stator vane platform <NUM> and the engine case structure <NUM> react different to the varying thermal environment.

For example, in the first stage of the high pressure turbine (HPT) (the designation "T1" referring to the first stage of the HPT) depicted in <FIG>, a transition of the outer stator vane platform <NUM> to the leading edge of the BOAS is defined to yield a smooth core flowpath at steady state conditions for maximum efficiency, typically at cruise conditions. However, relative motion or shifts between the T1 stator vane and the T1 BOAS segment due to the varying thermal environment can result in the BOAS leading edge portion moving radially outboard of the outer stator vane platform <NUM> (outer diameter platform) of the T1 stator vane, creating an outward step in the hot core flowpath from the outer stator vane platform <NUM> and BOAS (see, e.g., <FIG>). Relative motion or shifts may also result in the BOAS leading edge portion moving radially inboard of the outer stator vane platform <NUM> (outer diameter platform) of the T1 stator vane, making the BOAS leading edge portion project into the hot core flowpath (see, e.g., <FIG>). As noted previously, this results in higher heat load for the BOAS leading edge portion, shortening its life and/or compelling additional cooling.

However, according to various embodiments, the elongated transition portion 56b may improve gas flow transition across the leading edge wall 56a, and may prevent a stagnation region at the leading edge portion <NUM>. More particularly, various embodiments permit the transition from the upstream stator vane to the leading edge of the BOAS to be smoother and the leading edge portion less sensitive to being projected into the hot core flowpath air as a result of the relative movement/shifting of the BOAS segment and the surrounding static structure (e.g., the upstream stator vanes). As a result, various embodiments prolong BOAS life and/or tend to minimize cooling flow requirements for the BOAS segment/array, thereby maximizing turbine efficiency.

While various embodiments have been described to ease the transition between an upstream stator vane and an adjacent BOAS segment in a turbine stage, it is to be understood that various embodiments may be used to smooth the transition between adjacent non-rotating structures. As depicted in <FIG>, the proximate flowpath surface for the BOAS leading edge <NUM> of the main body <NUM> of BOAS segment <NUM> is the inner diameter flowpath surface <NUM>. Similar to elongated transition portion 56b of BOAS segment <NUM>, the stator vane <NUM> has an elongated transition portion 65b on outer stator vane platform <NUM> that transitions from a static combustor panel (not shown) at an outer platform leading edge <NUM> to the proximate flowpath surface 65c, the boundary between outer stator vane platform <NUM> and an airfoil <NUM> of stator vane <NUM>. Likewise, inner stator vane platform <NUM> includes an elongated transition portion 63b that extends from an inner platform leading edge <NUM> to a proximate flowpath surface 63c, the boundary between the elongated transition portion 63b and the airfoil <NUM> of stator vane <NUM>. Hence, the "proximate flowpath surface" may be an inner flowpath surface or an outer flowpath surface.

Claim 1:
A gas turbine engine component comprising a main body (<NUM>) having a leading edge (<NUM>) and a leading edge wall (56a) including an elongated transition portion (56b) extending between the leading edge (<NUM>) and an inner diameter flowpath surface (<NUM>) of the main body (<NUM>), wherein the elongated transition portion (56b) has an axial length (L) and a radial height and the axial length (L) is greater than the radial height, wherein the elongated transition portion (56b) is configured as an ellipse,
characterised in that:
the axial length (L) is defined from a first tangency point (<NUM>) defined between the leading edge (<NUM> and the leading edge wall (56a) to a second tangency point (<NUM>) defined between the leading edge wall (56a) and the inner diameter flowpath surface (<NUM>).