Airfoil with cooling circuit

An airfoil for a turbine engine having an engine component including an internal cooling circuit fluidly coupled to a plurality of passages within the outer wall of the engine component where cooling air moves from the internal cooling circuit to an outer surface of the engine component through the passages.

BACKGROUND OF THE INVENTION

Turbine engines, and particularly gas or combustion turbine engines, are rotary engines that extract energy from a flow of combusted gases passing through the engine onto a multitude of rotating turbine blades. Turbine engines have been used for land and nautical locomotion and power generation, but are most commonly used for aeronautical applications such as for aircraft, including helicopters. In aircraft, turbine engines are used for propulsion of the aircraft. In terrestrial applications, turbine engines are often used for power generation.

Turbine engines for aircraft are designed to operate at high temperatures to maximize engine efficiency, so cooling of certain engine components, such as the high pressure turbine and the low pressure turbine, can be beneficial. Typically, cooling is accomplished by ducting cooler air from the high and/or low pressure compressors to the engine components that require cooling. Temperatures in the high pressure turbine are around 1000° C. to 2000° C. and the cooling air from the compressor is around 500° C. to 700° C. While the compressor air is a high temperature, it is cooler relative to the turbine air, and can be used to cool the turbine.

Contemporary turbine blades generally include one or more interior cooling circuits for routing the cooling air through the blade to cool different portions of the blade. These cooling circuits can feed into smaller passages within the wall of the blade where cooling air can flow before exiting through film holes.

BRIEF DESCRIPTION OF THE INVENTION

An airfoil for a turbine engine, the airfoil comprising an outer wall having an outer surface and an inner surface bounding an interior space, the outer wall defining a pressure side and a suction side extending axially between a leading edge and a trailing edge and extending radially between a root and a tip, at least one coating applied to the outer surface, at least one cooling air supply circuit located within the interior, at least one wall cooling passage provided within at least a portion of an interior of the outer wall, at least one skin cooling circuit comprising at least one channel formed in the outer surface and at least one film hole passing through the coating to the channel, wherein the supply circuit is fluidly coupled to the wall cooling passage, which is fluidly coupled to the at least one channel to define a serial cooling air flow path via the supply circuit, wall cooling passage, and skin cooling circuit.

An engine component for a turbine engine, which generates a hot air flow, and provides a cooling fluid flow, the component comprising a wall separating the hot air flow from the cooling fluid flow and having a first surface along with the hot air flows in a hot flow path and a cooling surface facing the cooling fluid flow, at least one coating applied to the first surface, at least one wall cooling passage provided within at least a portion of an interior of the wall, at least one skin cooling circuit comprising at least one channel formed in the first surface and at least one film hole passing through the coating to the channel, wherein the cooling fluid flow is fluidly coupled to the wall cooling passage, which is fluidly coupled to the at least one channel to define a serial cooling air flow path.

A method of cooling an airfoil comprising passing a cooling air flow through an interior of an outer wall of an airfoil, then to a channel in an outer surface of the outer wall, and then to a film hole in a coating overlying the outer surface.

DESCRIPTION OF EMBODIMENTS OF THE INVENTION

The described embodiments of the present invention are directed to an engine component for an engine having a cooling circuit comprising skin cooling circuits coupled to wall cooling passages coupled to an internal passage or circuit within for example an airfoil to cool an outer surface of the airfoil. For purposes of illustration, the present disclosure will be described with respect to the turbine for an aircraft gas turbine engine. It will be understood, however, that the disclosure is not so limited and may have general applicability within an engine, including compressors, as well as in non-aircraft applications, such as other mobile applications and non-mobile industrial, commercial, and residential applications.

As used herein, the term “forward” or “upstream” refers to moving in a direction toward the engine inlet, or a component being relatively closer to the engine inlet as compared to another component. The term “aft” or “downstream” used in conjunction with “forward” or “upstream” refers to a direction toward the rear or outlet of the engine or being relatively closer to the engine outlet as compared to another component.

Additionally, as used herein, the terms “radial” or “radially” refer to a dimension extending between a center longitudinal axis of the engine and an outer engine circumference.

FIG. 1is a schematic cross-sectional diagram of a turbine engine10for an aircraft. The engine10has a generally longitudinally extending axis or centerline12extending forward14to aft16. The engine10includes, in downstream serial flow relationship, a fan section18including a fan20, a compressor section22including a booster or low pressure (LP) compressor24and a high pressure (HP) compressor26, a combustion section28including a combustor30, a turbine section32including a HP turbine34, and a LP turbine36, and an exhaust section38.

The fan section18includes a fan casing40surrounding the fan20. The fan20includes a plurality of fan blades42disposed radially about the centerline12. The HP compressor26, the combustor30, and the HP turbine34form a core44of the engine10, which generates a hot air flow. The core44is surrounded by core casing46, which can be coupled with the fan casing40.

A HP shaft or spool48disposed coaxially about the centerline12of the engine10drivingly connects the HP turbine34to the HP compressor26. A LP shaft or spool50, which is disposed coaxially about the centerline12of the engine10within the larger diameter annular HP spool48, drivingly connects the LP turbine36to the LP compressor24and fan20. The spools48,50are rotatable about the engine centerline and couple to a plurality of rotatable elements, which can collectively define a rotor51.

The LP compressor24and the HP compressor26respectively include a plurality of compressor stages52,54, in which a set of compressor blades56,58rotate relative to a corresponding set of static compressor vanes60,62(also called a nozzle) to compress or pressurize the stream of fluid passing through the stage. In a single compressor stage52,54, multiple compressor blades56,58can be provided in a ring and can extend radially outwardly relative to the centerline12, from a blade platform to a blade tip, while the corresponding static compressor vanes60,62are positioned upstream of and adjacent to the rotating blades56,58. It is noted that the number of blades, vanes, and compressor stages shown inFIG. 1were selected for illustrative purposes only, and that other numbers are possible.

The blades56,58for a stage of the compressor can be mounted to a disk61, which is mounted to the corresponding one of the HP and LP spools48,50, with each stage having its own disk61. The vanes60,62for a stage of the compressor can be mounted to the core casing46in a circumferential arrangement.

The HP turbine34and the LP turbine36respectively include a plurality of turbine stages64,66, in which a set of turbine blades68,70are rotated relative to a corresponding set of static turbine vanes72,74(also called a nozzle) to extract energy from the stream of fluid passing through the stage. In a single turbine stage64,66, multiple turbine blades68,70can be provided in a ring and can extend radially outwardly relative to the centerline12, from a blade platform to a blade tip, while the corresponding static turbine vanes72,74are positioned upstream of and adjacent to the rotating blades68,70. It is noted that the number of blades, vanes, and turbine stages shown inFIG. 1were selected for illustrative purposes only, and that other numbers are possible.

The blades68,70for a stage of the turbine can be mounted to a disk71, which is mounted to the corresponding one of the HP and LP spools48,50, with each stage having a dedicated disk71. The vanes72,74for a stage of the compressor can be mounted to the core casing46in a circumferential arrangement.

Complementary to the rotor portion, the stationary portions of the engine10, such as the static vanes60,62,72,74among the compressor and turbine section22,32are also referred to individually or collectively as a stator63. As such, the stator63can refer to the combination of non-rotating elements throughout the engine10.

In operation, the air flow exiting the fan section18is split such that a portion of the air flow is channeled into the LP compressor24, which then supplies pressurized air76to the HP compressor26, which further pressurizes the air. The pressurized air76from the HP compressor26is mixed with fuel in the combustor30and ignited, thereby generating combustion gases. Some work is extracted from these gases by the HP turbine34, which drives the HP compressor26. The combustion gases are discharged into the LP turbine36, which extracts additional work to drive the LP compressor24, and the exhaust gas is ultimately discharged from the engine10via the exhaust section38. The driving of the LP turbine36drives the LP spool50to rotate the fan20and the LP compressor24.

A portion of the pressurized air flow76can be drawn from the compressor section22as bleed air77. The bleed air77can be draw from the pressurized air flow76and provided to engine components requiring cooling. The temperature of pressurized air flow76entering the combustor30is significantly increased. As such, cooling provided by the bleed air77is necessary for operating of such engine components in the heightened temperature environments.

A remaining portion of the air flow78bypasses the LP compressor24and engine core44and exits the engine10through a stationary vane row, and more particularly an outlet guide vane assembly80, comprising a plurality of airfoil guide vanes82, at the fan exhaust side84. More specifically, a circumferential row of radially extending airfoil guide vanes82are utilized adjacent the fan section18to exert some directional control of the air flow78.

Some of the air supplied by the fan20can bypass the engine core44and be used for cooling of portions, especially hot portions, of the engine10, and/or used to cool or power other aspects of the aircraft. In the context of a turbine engine, the hot portions of the engine are normally downstream of the combustor30, especially the turbine section32, with the HP turbine34being the hottest portion as it is directly downstream of the combustion section28. Other sources of cooling fluid can be, but are not limited to, fluid discharged from the LP compressor24or the HP compressor26.

FIG. 2is a perspective view of an engine component in the form of one of the turbine blades68of the engine10fromFIG. 1. The turbine blade68includes a dovetail75and an airfoil79. The airfoil79extends radially between a root83and a tip81. The dovetail75further includes a platform85integral with the airfoil79at the root83, which helps to radially contain the turbine air flow. The dovetail75can be configured to mount to a turbine rotor disk on the engine10. The dovetail75comprises at least one inlet passage, exemplarily shown as a first inlet passage88, a second inlet passage90, and a third inlet passage92, each extending through the dovetail75to provide internal fluid communication with the airfoil79at a passage outlet94. It should be appreciated that the dovetail75is shown in cross-section, such that the inlet passages88,90,92are housed within the body of the dovetail75.

Turning toFIG. 3, the airfoil79, shown in cross-section, comprises an outer wall95bounding an interior space96having a concave-shaped pressure side98and a convex-shaped suction side100which are joined together to define an airfoil shape extending axially between a leading edge102and a trailing edge104. The blade68rotates in a direction such that the pressure side98follows the suction side100. Thus, as shown inFIG. 3, the airfoil79would rotate upward toward the top of the page.

Referring toFIG. 4, the interior space96can be divided into a plurality of internal cooling air supply circuits122,124,126which can be arranged in any formation within the interior space96and are dedicated to supply cooling air C to the interior space96. The supply circuits are fluidly coupled to at least one of the inlet passages88,90,92where internal fluid communication is provided to at least one of the supply circuits through the passage outlet94.

It should be appreciated that the respective geometries of each individual supply circuit within the airfoil79as shown is exemplary, and not meant to limit the airfoil to the number of supply circuits, their geometries, dimensions, or positions as shown. Additionally the supply circuits122,124,126can be fluidly coupled to each other to provide additional internal fluid communication between adjacent supply circuits. Further, while three supply circuits are shown, there can be any number of supply circuits from none, to one, to multiple, for example.

The outer wall95comprises an outer surface130and an inner surface132, which defines an interior134that is generally solid. At least one coating136is applied to the outer surface130where the coating136can include one or more layers comprising metallic, ceramic, or any other suitable material. The outer surface130and the at least one coating define a “skin” for the airfoil. The outer wall95, including the skin, separates a hot air flow H on a first surface128of the airfoil79from a cooling fluid flow comprising the cooling air C along a second surface129, and supplied to the cooling air circuits122,124,126. The coating136can be formed by various known methods such as spray, vapor deposition, and so forth, and also by additive manufacturing.

A plurality of film holes138can be provided through the surface of the outer wall95to provide cooling air C onto the exterior of the airfoil. A wall cooling circuit140and a skin cooling circuit142are provided with the airfoil and can be fluidly connected to one of the plurality of film holes138to cool the outer wall95and the skin130,136. It should be understood that film holes can be film cooling exits of any geometry, such as but not limited to holes, shaped holes, and slots.

Referring toFIG. 5, the details of the wall cooling circuit140and skin cooling circuit142will be described with respect to this schematic representation of a portion of the airfoil inFIG. 4. The wall cooling circuit140comprises one or more wall cooling passages144located within at least a portion of the interior134of the wall95bounded by the inner and outer surfaces130,132. Some of wall cooling passages144can be fluidly coupled in series. However, some or all of the wall cooling passages144can be fluidly independent from each other.

Some or all of the wall cooling passages144can be fluidly coupled to the supply circuit124through at least one internal hole152that can be formed as an aperture or slot. The cooling fluid flow C can flow from the supply circuit124into the internal hole152while flowing along the cooling surface129which can be any surface facing the cooling fluid flow C. It can be contemplated that at least some of the wall cooling passages144are fluidly coupled to different supply circuits wherein multiple wall cooling passages144circumscribe a portion of or the entire airfoil interior96and are fluidly coupled to supply circuits122and126.

The skin cooling circuit142comprises at least one channel150provided in the outer surface130and at least one film hole138passing through the coating136to the channel150. The skin cooling circuit142can be formed in the outer surface130or in the coating136or formed in a combination of both as illustrated. In some embodiments, a portion of the coating136shown could be part of the same substrate that forms the channels150, and then a coating added on top, enclosing the skin cooling circuit142with a non-coating material. It should be understood that the multiple channels and passages shown are exemplary and not meant to be limiting for example in shape, orientation, or size.

The at least one channel150can be multiple channels150, which are fluidly coupled to each other or fluidly separate from each other. The multiple channels150may be arranged in groups, which can be used to form sub-circuits within the skin cooling circuit142. The multiple channels150can vary in width and length as illustrated. It is further contemplated that multiple film holes138can pass through the coating to only one of the multiple channels150or to several or all of the multiple channels150.

As illustrated inFIG. 5, the wall cooling passages144are serially fluidly connected to each other as well as the channels150and the supply circuit124to define a serial cooling air flow path148via the supply circuit124, wall cooling passage144, and skin cooling circuit142. Other fluid connections are contemplated between the supply circuit124, wall cooling circuit140, and skin cooling circuit142. For example, any two or more of the supply circuits122,124,126, wall cooling circuit140, and skin cooling circuit144can be fluidly coupled, directly or indirectly, in any possible combination. For example some or all of the groups of skin cooling circuits142including each of the plurality of channels150can be directly or indirectly fluidly connected to the same or to a different one of multiple wall cooling passages144. Some or all of the wall cooling passages144can be fluidly connected to each other146and/or to the same or different supply circuit122,124,126. The skin cooling circuits142can directly connect141to the supply circuits or can be used to fluidly connect some of the wall cooling circuits140. The film holes138can directly connect to either or both of the wall cooling circuits140and the skin cooling circuits142. It can be contemplated that the channels150can be of the same or less dimensions as the wall cooling passages144, and in further embodiments are 50% or less of the wall cooling passages144.

Turning toFIG. 6, when cooling air is supplied to the supply circuit124in the form of the cooling fluid flow C, pressure differences between the supply circuit124and adjacent wall passages140cause the cooling air to move from the supply circuit124to the wall cooling passage140in a first flow direction SC defined by the supply circuit124where the cooling fluid flow C passes through the internal hole152.

A second flow direction WCP is dependent on the pressure differentials between the wall cooling passage144and the geometry of the adjacent skin cooling circuit124. Cooling air that comes into the wall cooling passage144from the supply circuit124will then move into one of the multiple channels150to define a third flow direction SCC which also depends on pressure differentials across the channels and individual film holes138.

While numerous flow paths are possible with the geometry of the configuration described herein, three possible flow paths are illustrated inFIGS. 7A, 7B, and 7C. These flow paths are for illustrative purposes only and are not meant to be limiting.

FIG. 7Aillustrates a flow path in which pressure differences between adjacent supply circuits124, wall cooling passages144, and channels150of corresponding skin cooling circuits142, produce first second and third flow directions SC, WCP, SCC where the second and third flows WCP, SCC are in the same direction creating a co-flow in which cooling air runs parallel with corresponding cooling air in adjacent passages and circuits while flowing in the same direction before exiting through the film holes138.

FIG. 7Billustrates a flow path in which pressure differences between adjacent supply circuits124, wall cooling passages144, and channels150of corresponding skin cooling circuits142, produce first second and third flow directions SC, WCP, SCC where the second and third flows WCC, SCC are in opposite directions creating a counter-flow in which cooling air runs parallel with corresponding cooling air in adjacent passages and circuits while flowing in the opposite direction as each other before exiting through the film holes138.

FIG. 7Cillustrates a flow path in which pressure differences between adjacent supply circuits124, wall cooling passages144, and channels150of corresponding skin cooling circuits142, produce first second and third flow directions SC, WCP, SCC where the second and third flows WCP, SCC are in both the same and opposite directions creating a split-flow in which cooling air runs parallel with corresponding cooling air in adjacent passages and circuits while flowing in both the same and opposite direction as each other before exiting through the film holes138.

It can be seen that a plurality of orientations can be contemplated regarding the geometry of the skin cooling circuit142relative to the wall cooling passages144. Some of the orientations are shown inFIGS. 8A, 8B, and 8C. While three different possible orientations are contemplated, it should be understood that they are for illustrative purposes only and are not meant to be limiting.FIG. 8Adepicts a relative geometry wherein channels250run perpendicular to wall passages244.FIG. 8Billustrates a relative geometry wherein channels350and wall passages344run parallel to each other. Furthermore,FIG. 8Cshows a relative orientation wherein channels450and wall passages444run at an angled orientation between the geometries depicted inFIGS. 8A and 6B.

Wall cooling passages144and skin cooling circuits142can co-flow, counter-flow, or flow at any straight or arcuate intermediate angles with respect to each other, as long as each layer of cooling is retained within its corresponding passage, aside from the fluid connections146between adjacent passages.

While illustrated on the suction side100, fluidly coupling the supply circuit122,124,126to the wall cooling passage144and the wall cooling passage144to the skin cooling circuit142and the skin cooling circuit142to the film holes138as described herein can be implemented at any location on the airfoil.

A method of cooling the airfoil79comprises first passing a cooling air flow through wall cooling passages144in the interior134of the outer wall95which can include multiple passages144. The method can further include passing cooling air between the multiple passages144. Cooling air is then passed to at least one of the channels150from at least some of the multiple passages144after which the cooling air passes on to at least one of the film holes138through the channels150. The method can include passing the cooling air through the interior134of the outer wall95in the different flow directions1,2,3described herein.

The passing of the cooling airflow through the skin cooling circuit can be a first direction SCC and the passing of the cooling airflow through the wall cooling circuit can be a second direction WCP wherein the first direction and second directions are the same, different, or opposite as described herein.

It can be contemplated that an entire engine component with multiple wall passages and skin cooling circuits can be cast as a single piece after which molding and reforming portions of the engine component can be implemented and then a coating applied. Additive manufacturing where a main component such as the supply circuit is cast and the additional components including the wall passages, fluidly connecting holes, and skin cooling circuits are added can also be contemplated.

Benefits to combining the wall cooling passage with skin cooling circuits include increasing thermal uniformity of the substrate structure. In the event that the more at risk skin cooling circuit fails, the underlying wall cooling passage can still provide thermal and structural integrity to the engine component.

In newer generation turbine cooling 30 to 50% less flow is utilized as compared to prior turbine cooling. Combining the wall cooling passage with skin cooling circuits enables 30% cooling flow reduction, decreases cost, and decreases specific fuel consumption.