Patent Application
20130232991
The present invention relates to turbine engines. In particular, the invention relates to internal cooling channel pedestals of an airfoil for a turbine engine.
A turbine engine employs a variety of airfoils to extract energy from a flow of combustion gases to perform useful work. Some airfoils, such as, for example, stator vanes and rotor blades, operate downstream of the combustion gases and must survive in a high-temperature environment. Often, airfoils exposed to high temperatures are hollow, having internal cooling channels that direct a flow of cooling air through the airfoil to remove heat and prolong the useful life of the airfoil. A source of cooling air is typically taken from a flow of compressed air produced upstream of the stator vanes and rotor blades. Some of the energy extracted from the flow of combustion gases must be used to provide the compressed air, thus reducing the energy available to do useful work and reducing an overall efficiency of the turbine engine.
Internal cooling channels are designed to provide efficient transfer of heat between the airfoils and the flow of cooling air within. As heat transfer efficiency improves, less cooling air is necessary to adequately cool the airfoils. Internal cooling channels typically include structures to improve heat transfer efficiency including, for example, pedestals (also known as pin fins). Pedestals link opposing sides of such airfoils (pressure side and suction side) to improve heat transfer by increasing both the area for heat transfer and the turbulence of the cooling air flow. The improved heat transfer efficiency results in improved overall turbine engine efficiency.
While the use of hollow airfoils provides for a flow of cooling air to extend the useful life of the airfoils, hollow blades are not as mechanically strong as solid blades. Improvements to the mechanical strength of hollow airfoils are needed to further extend their useful life.
An embodiment of the present invention is an airfoil for a turbine engine, the airfoil including a first side wall, a second side wall spaced apart from the first side wall, and an internal cooling channel formed between the first side wall and the second side wall. The internal cooling channel includes at least one pedestal having a first pedestal end connected to the first side wall and a second pedestal end connected to the second side wall. The internal cooling channel also includes a first fillet and a second fillet. The first fillet is disposed around the periphery of the first pedestal end between the first side wall and the first pedestal end. The second fillet is disposed around the periphery of the second pedestal end between the second side wall and the second pedestal end. At least one of the first fillet and the second fillet includes a profile that is non-uniform around the periphery of the corresponding pedestal end.
The present invention provides for greater mechanical strength and durability of pedestals in an internal cooling channel within an airfoil by employing fillets around the periphery of pedestal ends where the pedestal ends connect to airfoil walls. The fillets each have a profile that is non-uniform around the periphery of the corresponding pedestal end. While larger fillets provide greater mechanical strength, larger fillets also obstruct the flow of cooling air through the internal cooling channel, thereby reducing the heat transfer efficiency gains provided by the pedestals. The non-uniform fillet of the present invention is smaller around most of the periphery of the pedestal end to reduce the obstruction of cooling air flow and larger only at those points likely to experience the highest levels of mechanical stress and serve as initiation points for pedestal connection failure.
As illustrated in
In operation, air enters compressor section 14 through fan 12. The air is compressed by the rotation of compressor section 14 driven by high-pressure rotor 20. The compressed air from compressor section 14 is divided, with a portion going to combustor section 18, and a portion employed for cooling airfoils, such as rotor blades 24 and stator vanes 26, as described below. Compressed air and fuel are mixed an ignited in combustor section 16 to produce high-temperature, high-pressure combustion gases. The combustion gases exit combustor section 16 into turbine section 18 Stator vanes 26 properly align the flow of the combustion gases for an efficient attack angle on rotor blades 24. Because rotor blades 24 include an airfoil section, the flow of combustion gases past rotor blades 24 drives rotation of both high-pressure rotor 20 and low-pressure rotor 22. High-pressure rotor 20 drives compressor section 14, as noted above, and low-pressure rotor 22 drives fan 16 to produce thrust from gas turbine engine 10. Although embodiments of the present invention are illustrated for a turbofan gas turbine engine for aviation use, it is understood that the present invention applies to other aviation gas turbine engines and to industrial gas turbine engines as well.
Rotor blades 24 spin at relatively high revolutions per minute, resulting in significant mechanical stress on rotor blades 24. In addition, as rotor blades 24 spin past stator vanes 26, they experience a varying flow of combustion gases which causes a change in force experienced by rotor blades 24. A sequence of changing forces experienced by rotor blades 24 as they spin past stator vanes 26 causes a vibratory motion in rotor blades 24 causing warping, or twisting of the airfoil section of rotor blades 24 about each of their respective vertical axes. This warping stress presents a particular challenge to mechanical structures within the airfoil section. As described below, rotor blades 24 embodying the present invention are strengthened to meet this challenge.
As mentioned above, airfoils operating downstream of combustor section 16, such as stator vanes 26 and rotor blades 24, operate in a high-temperature environment. Often, airfoils exposed to high temperatures are hollow, having internal cooling channels that direct a flow of cooling air through the airfoil to remove heat and prolong the useful life of the airfoil.
Platform 132 connects one end of airfoil section 134 to root section 130. Thus, leading edge 136, trailing edge 138, suction side wall 140, and pressure side wall 142 extend from platform 132. Tip 144 closes off the other end of airfoil section 134. Suction side wall 140 and pressure side wall 142 connect leading edge 136 and trailing edge 138. Film cooling holes 146 are arranged over the surface of airfoil section 134 to provide a layer of cool air proximate the surface of airfoil section 134 to protect it from high-temperature combustion gases. Trailing edge slots 148 are arranged along trailing edge 138 to provide an exit for air circulating within airfoil section 134, as described below in reference to
Considering
In operation, rotor blade 24 is exposed not only to high-temperature combustion gases, but to extreme mechanical stresses, including the warping stress experienced by airfoil section 134 described above. Warping stress experienced by airfoil section 134 creates a mechanical stress at locations where pedestal 154 connects to suction side wall 140 and where pedestal 154 connects to pressure side wall 142. Such mechanical stresses can result in mechanical failure of one of the pedestal connections. The present invention employs fillets around the periphery of pedestal 154, between first end 156 and suction side wall 140 and between second end 158 and pressure side wall 142. Fillets spread the stress at the pedestal connections over a larger area, reducing the level of stress at any particular location to prevent mechanical failure. Larger fillets spread the stress over a larger area, protecting against a higher level of warping stress. However, larger fillets obstruct serpentine flow channel 152, and the flow of cooling air, thereby reducing the heat transfer efficiency gains provided by pedestals 154. Thus, determining the proper fillet size involves a trade off between mechanical durability and heat transfer efficiency. The present invention overcomes this problem with a fillet that is smaller around most of the periphery of the pedestal end and larger only at those points likely to experience the highest levels of mechanical stress and serve as initiation points for pedestal connection failure.
In this embodiment, first fillet 160 and second fillet 162 are concave and their respective profiles at any point around the periphery of the corresponding pedestal end may be described by a simple curve, that is, described by a single radius of curvature at that point. However, it is understood that other profiles are encompassed by the present invention, including compound curves, as described below in reference to
The side cross-sectional view of
As shown in
The side cross-sectional view of
In embodiments described above, first fillets and second fillets are illustrated as mirror images on either end of the pedestal, such as first fillet 160 and second fillet 162 on either end of pedestal 154 as described above in reference to
The present invention has been described in detail with respect to rotor blades. However, it is understood that the present invention encompasses embodiments in which the airfoil section is a stator vane, such as stator vane 26. Although stator vanes are not subject to stresses as severe as rotor blades, stator vanes are nonetheless subject to warping stresses due to reaction forces from their proximity to spinning rotor blades.
For simplicity in illustration and to avoid unnecessary repetition, many of the embodiments are described above with a larger portion of a non-uniform fillet nearer a leading edge of an airfoil. However, it is understood that the present invention also encompasses embodiments where a larger portion of a non-uniform fillet is nearer a trailing edge of an airfoil. Similarly, use of a serpentine cooling channel leading to a trailing edge of an airfoil, with a pedestal array near the trailing edge is merely exemplary. It is understood that the present invention encompasses embodiments where the internal cooling channel is of other shapes and varieties, including, for example, multi-walled internal cooling channels where the side walls to which pedestal ends attach are not a pressure side wall or a suction side wall. The present invention also encompasses embodiments where pedestals are not near the trailing edge of an airfoil.
A method for providing enhanced gas turbine engine airfoil durability begins with introducing cooling air into an internal cooling channel within the airfoil. The cooling air flows through the internal cooling channel past pedestals connected to walls of the airfoil. The internal cooling channel includes fillets at pedestal ends, at least some of the fillets including a profile that is non-uniform around the periphery of the corresponding pedestal end. Finally, cooling air is exhausted through the trailing edge cooling slot.
The present invention provides for greater mechanical strength and durability of pedestals in an internal cooling channel within an airfoil by employing fillets around the periphery of pedestal ends where the pedestal ends connect to airfoil walls. The fillets each have a profile that is non-uniform around the periphery of the corresponding pedestal end. The non-uniform fillet of the present invention is smaller around most of the periphery of the pedestal end to reduce the obstruction of cooling air flow and larger only at those points likely to experience the highest levels of mechanical stress and serve as initiation points for pedestal connection failure.
While the invention has been described with reference to an exemplary embodiment(s), it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the invention. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the invention without departing from the essential scope thereof. Therefore, it is intended that the invention not be limited to the particular embodiment(s) disclosed, but that the invention will include all embodiments falling within the scope of the appended claims.
The following are non-exclusive descriptions of possible embodiments of the present invention.
An airfoil for a turbine engine can include a first side wall; a second side wall spaced apart from the first side wall; and an internal cooling channel formed between the first side wall and the second side wall, the internal cooling channel including at least one pedestal having a first pedestal end connected to the first side wall and a second pedestal end connected to the second side wall; a first fillet disposed around the periphery of the first pedestal end between the first side wall and the first pedestal end; and a second fillet disposed around the periphery of the second pedestal end between the second side wall and the second pedestal end; wherein at least one of the first fillet and the second fillet includes a profile that is non-uniform around the periphery of the corresponding pedestal end.
The airfoil of the preceding paragraph can optionally include, additionally and/or alternatively, any one or more of the following features, configurations and/or additional components:
the airfoil is one of a turbine rotor blade and a turbine stator vane;
the pedestal is one of a cylinder and an elliptic cylinder;
the airfoil further includes a leading edge; a trailing edge; a pressure side wall connecting the leading edge and the trailing edge; and a suction side wall spaced apart from the pressure side wall, the suction side wall connecting the leading edge and the trailing edge; wherein the pressure side wall is the first side wall and the suction side wall is the second side wall;
the profile is a simple curve described at any point around the periphery of the corresponding pedestal end by a radius of curvature at a point; the profile at a first point includes a first local maximum value of the radius of curvature; the first point being a point around the periphery nearest the leading edge;
the profile at a second point includes a second local maximum value of the radius of curvature, the second point being a point around the periphery nearest the trailing edge;
the profile is a compound curve described at any point by a first radius of curvature describing a first portion of the profile at that point and a second radius of curvature describing a second portion of the profile at that point, each radius having a different center point; the first portion being closer to the corresponding one of the pressure side wall and the suction side wall than the second portion; the profile at a first point includes a first local maximum value of the first radius of curvature; the first point being a point around the periphery nearest the leading edge;
the profile is a simple curve described at any point by a radius of curvature at that point; the profile at a first point includes a first local maximum value of the radius of curvature; the first point between a second point around the periphery nearest the leading edge, and a third point around the periphery nearest the trailing edge;
the first point is closer to the second point than to the third point;
the airfoil further includes a platform from which the leading edge, trailing edge, pressure side wall, and suction side wall extend; wherein the first point is closer to the platform than either of the second point or the third point; and/or
the airfoil further includes a platform from which the leading edge, trailing edge, pressure side wall, and suction side wall extend; wherein the first point is farther from the platform than either of the second point or the third point.
A gas turbine engine can include a compressor section; a combustor section; and a turbine; the turbine including a plurality of airfoils, at least one of the plurality of airfoils including a first side wall; a second side wall spaced apart from the first side wall; and an internal cooling channel formed between the first side wall and the second side wall, the internal cooling channel including at least one pedestal having a first pedestal end connected to the first side wall and a second pedestal end connected to the second side wall; a first fillet disposed around the periphery of the first pedestal end between the first side wall and the first pedestal end; and a second fillet disposed around the periphery of the second pedestal end between the second side wall and the second pedestal end; wherein at least one of the first fillet and the second fillet includes a profile that is non-uniform around the periphery of the corresponding pedestal end.
The engine of the preceding paragraph can optionally include, additionally and/or alternatively, any one or more of the following features, configurations and/or additional components:
wherein the at least one of the plurality of airfoils is one of a rotor blade and a stator vane;
wherein the pedestal is one of a cylinder and an elliptic cylinder;
the least one of the plurality of airfoils further includes a leading edge; a trailing edge; a pressure side wall connecting the leading edge and the trailing edge; and a suction side wall spaced apart from the pressure side wall, the suction side wall connecting the leading edge and the trailing edge; wherein the pressure side wall is the first side wall and the suction side wall is the second side wall;
the profile is a simple curve described at any point around the periphery of the corresponding pedestal end by a radius of curvature at that point; the profile at a first point includes a first local maximum value of the radius of curvature; the first point being a point around the periphery nearest the leading edge;
the profile at a second point includes a second local maximum value of the radius of curvature, the second point being a point around the periphery nearest the trailing edge;
the profile is a compound curve described at any point by a first radius of curvature describing a first portion of the profile at that point and a second radius of curvature describing a second portion of the profile at that point, each radius having a different center point; the first portion being closer to the corresponding one of the pressure side wall and the suction side wall than the second portion; the profile at a first point includes a first local maximum value of the first radius of curvature; the first point being a point around the periphery nearest the leading edge;
the profile is a simple curve described at any point by a radius of curvature at that point; the profile at a first point includes a first local maximum value of the radius of curvature; the first point between a second point around the periphery nearest the leading edge, and a third point around the periphery nearest the trailing edge;
the first point is closer to the second point than to the third point;
the engine further includes a platform from which the leading edge, trailing edge, pressure side wall, and suction side wall extend; wherein the first point is closer to the platform than either of the second point or the third point; and/or
the engine further includes a platform from which the leading edge, trailing edge, pressure side wall, and suction side wall extend; wherein the first point is farther from the platform than either of the second point or the third point.
A method for providing enhanced gas turbine engine airfoil durability, the method includes introducing cooling air into an internal cooling channel within the airfoil; flowing the cooling air through the internal cooling channel past pedestals connected to walls of the airfoil; the internal cooling channel including fillets at pedestal ends, at least some of the fillets including a profile that is non-uniform around the periphery of the corresponding pedestal end; and exhausting cooling air through trailing edge cooling slots.
