US20160146019A1 - Cooling channel for airfoil with tapered pocket - Google Patents

Abstract

The present invention includes systems and methods for providing cooling channels located within walls of a turbine airfoil. These cooling channels include micro-circuits that taper in various directions along the length and width of the airfoil. In addition, these cooling channels have a variety of shapes and areas to facilitate convective heat transfer between the surrounding air and the airfoil.

Description

CROSS-REFERENCE TO RELATED APPLICATIONS
• This Application claims the benefit of U.S. Provisional Application No. 62/084,810, filed Nov. 26, 2014, and titled “GAS TURBINE AIRFOIL WITH TAPERED AIRFLOW MICRO CIRCUITS FOR IMPROVED COOLING,” which is incorporated herein by reference in its entirety. This application is also related by subject matter to concurrently filed U.S. patent application Ser. No. (not yet assigned; Attorney Docket No. PSSF.241854), filed Nov. 24, 2015, and titled “LEADING EDGE COOLING CHANNEL FOR AIRFOIL,” and concurrently filed U.S. patent application Ser. No. (not yet assigned; Attorney Docket No. PSSF.241855), filed Nov. 24, 2015, and titled “TAPERED COOLING CHANNEL FOR AIRFOIL.” The teachings of each of these concurrently filed applications are also incorporated herein by reference in their entirety.
TECHNICAL FIELD
• The present invention relates to turbine airfoils, and more particularly, to cooling circuits incorporated into turbine airfoils.
BACKGROUND OF THE INVENTION
• A typical gas turbine engine is comprised of three main sections: a compressor section, a combustor section, and a turbine section. When in a standard operating cycle, the compressor section is used to pressurize air supplied to the combustor section. In the combustor section, a fuel is mixed with the pressurized air from the compressor section and is ignited in order to generate high temperature and high velocity combustion gases. These combustion gases then flow into a multiple stage turbine, where the high temperature gas flows through alternating rows of rotating and stationary gas turbine airfoils. The rows of stationary vanes are typically used to redirect the flow of combustion gases onto a subsequent stage of rotating blades. The turbine section is coupled to the compressor section along a common axial shaft, such that the turbine section drives the compressor section.
• The air and hot combustion gases are directed through a turbine section by turbine blades and vanes. These blades and vanes are subject to extremely high operating temperatures, often exceeding the material capability from which the blades and vanes are made. Extreme temperatures can also cause thermal growth in the components, thermal stresses, and can lead to durability shortfall. In order to lower the effective operating temperature, the blades and vanes are cooled, often with air or steam. However, the cooling must occur in an effective way so as to use the cooling fluid efficiently. As a result, an improved cooling design for airfoils in gas turbines that addresses these issues, among others, is needed.
BRIEF SUMMARY OF THE INVENTION
• In brief, and at a high level, the subject matter of this application relates generally to cooling passages, channels, and chambers incorporated into gas turbine airfoils. A gas turbine airfoil is comprised of an airfoil wall that includes an inner surface and an outer surface, and that forms an airfoil chamber that is at least partially enclosed by the airfoil wall. Embodiments provide for airfoil passages and pockets that are formed in various locations, directions, and configurations in the airfoil wall for improved cooling of the airfoil. The airfoil passages allow for cooling fluid or air to pass through the airfoil wall and airfoil chamber, cooling the airfoil during operation of the gas turbine.
• In a first embodiment of the invention, an airfoil for a gas turbine having a leading edge and a trailing edge is provided. The airfoil comprises an airfoil wall having an inner surface and an outer surface, the airfoil wall forming an airfoil chamber at least partially enclosed by the airfoil wall, and a plurality of pockets within the airfoil wall, each of the plurality of pockets comprising an inner pocket wall and an outer pocket wall, at least one first opening in the inner surface positioned at a first distance away from the leading edge, the first opening providing fluid communication between the airfoil chamber and the pocket, and at least one second opening in the outer surface positioned at a second distance away from the leading edge, the second opening providing fluid communication between an outside of the airfoil and the pocket. A distance between the inner pocket wall and the outer pocket wall is larger proximate the leading edge and smaller proximate the trailing edge.
• In another embodiment of the present invention, a gas turbine assembly is provided. The assembly comprises an airfoil having a leading edge and a trailing edge, the airfoil comprising an airfoil wall having an inner surface and an outer surface, the airfoil wall forming an airfoil chamber at least partially enclosed by the airfoil wall, and a plurality of pockets within the airfoil wall, each of the plurality of pockets comprising an inner pocket wall and an outer pocket wall, at least one first opening in the inner surface positioned at a first distance away from the leading edge, the first opening providing fluid communication between the airfoil chamber and the pocket, and at least one second opening in the outer surface positioned at a second distance away from the leading edge, the second opening providing fluid communication between an outside of the airfoil and the respective pocket. A distance between the inner pocket wall and the outer pocket wall is larger proximate the leading edge and smaller proximate the trailing edge, and a ratio of cross-sectional area of the inner pocket wall to the outer pocket wall is between 1.1:1 and 10:1.
• In yet another embodiment of the present invention, a method of manufacturing airfoils is provided. The method comprises providing an airfoil having a leading edge and a trailing edge, the airfoil having an airfoil wall having an inner surface and an outer surface, the airfoil wall forming an airfoil chamber at least partially enclosed by the airfoil wall, and forming a plurality of pockets within the airfoil wall, each of the plurality of pockets comprising an inner pocket wall and an outer pocket wall, a first opening in the inner surface positioned at a first distance away from the leading edge, the first opening providing fluid communication between the airfoil chamber and the pocket, and a second opening in the outer surface positioned at a second distance away from the leading edge, the second opening providing fluid communication between an outside of the airfoil and the pocket. A distance between the inner pocket wall and the outer pocket wall is larger proximate a leading edge of the airfoil and smaller proximate a trailing edge of the airfoil.
• The cooling circuits, channels, passages, and/or micro-circuits described in this disclosure are discussed frequently in the context of gas turbine airfoils, but may be used in any type of airfoil structure. Additionally, cooling fluid, gas, air, and/or airflow may be used interchangeably in this disclosure, and refer to any cooling medium that can be sent through an airfoil to provide heat transfer and cooling of the airfoil.
BRIEF DESCRIPTION OF THE SEVERAL VIEWS OF THE DRAWING
• The present invention is described in detail below with reference to the attached drawing figures, wherein:
• FIG. 1A is a perspective view of a prior art gas turbine airfoil;
• FIG. 1B is a perspective view of a prior art gas turbine vane;
• FIG. 2 is a cross-sectional view of the airfoil shown inFIG. 1A;
• FIG. 3A is an angled, perspective, cross-sectional view of an airfoil with cooling channels, in accordance with an embodiment of the present invention;
• FIG. 3B is a cross-sectional view of the airfoil shown inFIG. 3A, in accordance with an embodiment of the present invention;
• FIG. 3C is a partial, cross-sectional, perspective view of a cooling pocket of the airfoil shown inFIGS. 3A and 3B, in accordance with an embodiment of the present invention;
• FIG. 4A is a cross-sectional view of an airfoil with a first configuration of cooling channels, in accordance with an embodiment of the present invention;
• FIG. 4B is a partial, perspective, cross-sectional view of the airfoil shown inFIG. 4A, in accordance with an embodiment of the present invention;
• FIG. 4C is a perspective view of a radially tapering airfoil passage which can be formed into an airfoil wall, in accordance with an embodiment of the present invention;
• FIG. 4D is a cross-sectional view of the airfoil passage shown inFIG. 4C incorporated into an airfoil wall and including a flow turbulator, in accordance with an embodiment of the present invention;
• FIG. 5A is a perspective view of an airfoil having multiple cooling channels, in accordance with an embodiment of the present invention;
• FIG. 5B is a cross-sectional, elevation view of the airfoil shown inFIG. 5A, in accordance with an embodiment of the present invention;
• FIG. 6 is a partial, angled, perspective view of cooling channels incorporated into a leading edge of an airfoil, in accordance with an embodiment of the present invention;
• FIGS. 7A, 7B, and 7C are cut views of various cooling channel designs which can be incorporated into an airfoil, in accordance with embodiments of the present invention;
• FIG. 8 is a cut view of an alternate cooling channel design, in accordance with an embodiment of the present invention;
• FIGS. 9A and 9B are cut views of alternate cooling channel designs, in accordance with embodiments of the present invention;
• FIGS. 10A, 10B, and 10C are cut views of alternate cooling channel designs, in accordance with embodiments of the present invention;
• FIG. 11 is a block diagram of an exemplary method of manufacturing gas turbine airfoils, in accordance with an embodiment of the present invention;
• FIG. 12 is a block diagram of an exemplary method of manufacturing gas turbine airfoils, in accordance with an embodiment of the present invention; and
• FIG. 13 is a block diagram of an exemplary method of manufacturing gas turbine airfoils, in accordance with an embodiment of the present invention.
DETAILED DESCRIPTION OF THE INVENTION
• At a high level, the subject matter of this application generally relates to an airfoil for a gas turbine that includes cooling circuits integrated in various configurations. The airfoil may generally include an airfoil wall with an inner surface and an outer surface that at least partially encloses an airfoil chamber. Cooling circuits may be formed in various locations in the airfoil wall, to provide enhanced heat transfer from the airfoil when the gas turbine is in operation and cooling fluid or gas is passing through the cooling circuits. For turbine hardware operating in harsh environments, the use of this airfoil cooling technology is fully contemplated to be adapted to additional components such as outer and inner diameter platforms, blade outer or inner air shields, or alternative high temperature turbine components.
• Referring now toFIG. 1A, agas turbine blade100 is provided. Theturbine blade100 comprises a bottom portion commonly referred to as aroot102, which may be coupled to a rotor disk (not shown). It is understood that the root may be completely integrated into the rotor disk, such that the root does not extend into the flow path. Extending in an upward radial, typically perpendicular to the rotor central axis, direction from theroot102 is theneck103. Theneck103 may primarily be used as a transitional piece between theroot102 and thegas turbine airfoil104.
• Thegas turbine airfoil104 is comprised of four distinct portions. The first portion of theairfoil104 that comes into contact with pressurized gas flow is referred to as theleading edge106, which is opposed by the last portion of the airfoil to come in contact with the gas flow, defined as the trailingedge108. Theleading edge106 faces the turbine compressor section (not shown), or turbine inlet, along the rotor center axis. This direction is referred to as the axial direction. When pressurized airflow impedes upon theleading edge106, the airflow splits into two separate streams of air with different relative pressures. Connecting theleading edge106 and the trailingedge108 are two radially extending walls, which are defined based on the relative pressures impeding on the walls. The concave surface seen inFIG. 1A is defined to be apressure side wall110. The concave geometry of this surface generates a higher local pressure along the length of thepressure side wall110. Opposing thepressure side wall110 is asuction side wall112. Thesuction side wall112 has a convex geometry, which generates a lower local pressure along the length of thesuction side wall112.
• The pressure differential created between thepressure side wall110 and thesuction side wall112 creates an upward lifting force along the cross-section of thegas turbine airfoil104. The cross-section of thegas turbine airfoil104 can be seen in greater detail in FIG.2. This lifting force actuates the rotational motion of the rotor disk. The rotor disk may be coupled to a compressor and a generator via a shaft (not shown) for the purposes of generating electricity. The uppermost portion ofFIG. 1A shows atip shroud114 containing afirst surface116 that is populated withknife edges118 that extend radially outward from thefirst surface116. Located between the knife edges118 are recessedpockets120.
• Avane assembly150 of the prior art is shown inFIG. 1B, and comprises aninner platform151,inner rail152,outer platform153, andvane airfoils154 extending betweeninner platform151 andouter platform153. While theinner rail152 serves as a means to seal the rim cavity region from leakage of the cooling air into the hot gas path instead of passing to the designated vanes,inner rail152 also stiffensinner platform151.Inner rail152 may be located proximate the plenum of cooling air and therefore operates at approximately the temperature of the cooling air.
• FIG. 2 is a cross-sectional view of a prior art cooling design for a gas turbine airfoil.FIG. 2 is cross-sectional for the purposes of showingcooling passages202 and203. Gas turbine airfoils may operate in an environment where temperatures exceed the melting point of the materials used to construct the airfoil. Therefore, coolingpassages202 and203 are provided as a way to decrease the temperature of the airfoil during operation by flowing cooling air through the cooling passages of the airfoil.
• Traditionally, air cooled turbine airfoils are produced by a machining process or an investment casting process by forming a wax body of the turbine airfoil, providing an outer shell about the wax part, and then melting the wax to leave a mold for the liquid metal. Then, liquid metal is poured into the mold to fill the void left by the wax. Often-times the wax also contains a ceramic core to establish cooling channels within the metal turbine airfoils. Once the liquid metal cools and solidifies, the shell is removed and the ceramic core is chemically leached out of the now solid metal turbine airfoil, resulting in a hollow turbine airfoil. These traditional casting methods have limits as to the geometry that can be cast. New developments in additive manufacturing have occurred which can expand the capabilities beyond traditional investment casting techniques.
• The turbine airfoils ofFIGS. 1A, 1B, and 2 are known to be manufactured using standard metallurgy techniques, such as investment casting. However, the geometries that can be created using traditional manufacturing technique are limited. Internal geometrical shapes, as well as small geometrical intricacies, are generally not suitable for die casting. Advances in the field of additive manufacturing, have been adopted for the manufacturing of intricacies that were previously unattainable. The embodiments of the present invention may be created using an additive manufacturing process. An example of an additive manufacturing process is selective laser melting, known more commonly in the manufacturing field as SLM. Although SLM is widely considered a common additive manufacturing process, the embodiments described herein can be manufactured with any additive manufacturing process, such as selective laser sintering (SLS) or direct metal laser sintering (DMLS) or an alternative additive manufacturing method. The SLM processes described herein are intended to be non-limiting and exemplary.
• FIGS. 3A and 3B are cross-sectional perspective views of an exemplarygas turbine airfoil300 incorporating various cooling channels, in accordance with an embodiment of the present invention. Theairfoil300 includes anairfoil wall301 having aninner surface303 and anouter surface305. Theairfoil wall301 at least partially encloses anairfoil chamber307 within theairfoil wall301. Theairfoil wall301 as a whole comprises aleading edge302, a trailingedge304, apressure side wall306, and asuction side wall308. Positioned within thepressure side wall306 arepockets310 and312.Pockets314 and316 are positioned within thesuction side wall308. Thesepockets310,312,314, and316 have been introduced into theairfoil wall301 of thegas turbine airfoil300 for the purpose of increasing active cooling within theairfoil300 by allowing cooling fluid or gas to pass through interior portions of theairfoil wall301 to carry heat away from theairfoil300 during operation of an associated gas turbine to which theairfoil300 is coupled.
• Additionally, thepocket sections310,312,314, and316 (which are shown by the spaces within the airfoil wall301) may be manufactured using an additive manufacturing process, as previously discussed. As shown inFIGS. 3A and 3B, pockets310,312,314, and316 each extend within theairfoil wall301, and each include afirst opening318, which may be one of a plurality offirst openings318, referred to hereinafter as thefirst opening318 for simplicity but intended to be non-limiting, (which may be a cooling fluid inlet) on theinner surface303, and asecond opening320, which may be one of a plurality ofsecond openings320, referred to hereinafter as thesecond opening320 for simplicity but intended to be non-limiting (which may be a cooling fluid outlet) on theouter surface305. Theseopenings318,320 are provided and paired for each of thepockets310,312,314, and316. Thefirst opening318 of each of thepockets310,312,314 and316 provides fluid communication between theairfoil chamber307 and therespective pocket310,312,314 or316, and thesecond opening320 provides fluid communication between therespective pockets310,312,314 or316 and an outside environment of theairfoil300. Theseopenings318,320 feed and exhaust theinterior pockets310,312,314, and316 of the airfoil shown inFIGS. 3A-3C.
• Included within each of thepockets310,312,314, and316 of theairfoil wall301 are a plurality ofpedestals322, which extend between aninner pocket wall324 and anouter pocket wall326 of each of thepocket310,312,314, and316. Thepockets310,312,314, and316 may each include one or more flow turbulators (not shown), which may be extruded portions of thepocket310,312,314, or316 that promote turbulent mixing of cooling fluid or gas, to provide further sidewall cooling. These can be implemented or included as various different structures or extrusions, simply to provide mixing of cooling fluid traveling between the respectivefirst opening318 and respectivesecond opening320 within thepockets310,312,314, and316. Turbulation may alternatively be achieved by manufacturing pockets having a rough surface. The topography of a surface with roughness is complex and there is no single definitive measure of roughness. A widely used basic perimeter is “equivalent roughness” (Ra), defined as the arithmetic average of the absolute values of the measured profile height deviations of the surface from the surface profile centerline within a given sampling length. Typical values of Ra for turbomachinery components are 125 micro-inches for material as cast and 25 micro-inches for polished components. In the disclosed embodiments, the pocket heat transfer coefficient may be additionally modified by tailoring the surface roughness to achieve an equivalent roughness measured value of at least 400 Ra.
• Thepockets310,312,314, and316 are included in an airfoil side wall and taper in an area generally along the axial direction from theleading edge302 to the trailingedge304. The taper is a reduction in cross-sectional area between thefirst opening318 andsecond opening320 of eachrespective pocket310,312,314, and316. The ratio of cross-sectional area difference between thefirst opening318 and thesecond opening320 of each of thepockets310,312,314, and316 may vary between 1.1:1 and 10:1, in order to accelerate the flow of cooling fluid traveling between thefirst opening318 and thesecond opening320 within each of therespective pockets310,312,314, and316. This results in a balance between the internal heat pick-up and heat transfer coefficient. In other words, as more heat is removed from theairfoil300 through passage of the cooling fluid or gas through therespective pockets310,312,314, and316, the cooling fluid or gas becomes hotter and able to absorb less heat from theairfoil wall301, and the acceleration of the cooling fluid or gas within therespective pockets310,312,314, and316 allows the cooling fluid or gas to at least partially maintain the desired heat transfer coefficient through thepockets310,312,314, and316. In this embodiment, the reduction in cross-sectional area tapers in an axial direction, as the reduction in cross-sectional area occurs in the direction of cooling passage flow between thefirst opening318 andsecond opening320 generally along the axis of the rotor disk (not shown).
• InFIGS. 3A and 3B, the distance between theinner pocket wall324 andouter pocket wall326 may be larger proximate theleading edge302 of theairfoil300 and smaller proximate the trailingedge304 of theairfoil300. This internal passage differentiation may be further characterized by a ratio of pocket length (axial or radial) to airfoil wall width. The airfoil wall width is defined as the thickness between theinner surface303 and theouter surface305 of theairfoil300. The pocket length, fully enclosed within theairfoil wall301 in a generally axial direction, to airfoil wall width may be a minimum ratio of 1:1 to a maximum ratio dependent upon an airfoil span between theleading edge302 and the trailingedge304 of theairfoil300. This minimum ratio may also be described as the pocket length to pocket width, defined as distance between theinner pocket wall324 and theouter pocket wall326 measured at thefirst opening318, as a minimum ratio of 3:1.
• Additionally, it is contemplated herein that each of the plurality ofpedestals322 inFIGS. 3A, 3B, and most clearly shown inFIG. 3C, may have a circular, triangular, square, ovular, or rectangular cross-sectional shape, among other shapes. Further, each of the plurality ofpedestals322 may have a non-uniform or varying cross-sectional area, for the purposes of creating optimal air flow characteristics within eachpocket310,312,314, and316.
• Also, inFIGS. 3A and 3B,pocket sections310,312,314 and316 may be arrayed in a linear or non-linear pattern within theairfoil wall301, or rather, not aligned linearly along theairfoil wall301. Further, the shape of theinner pocket wall324 and theouter pocket wall326 may be aligned substantially parallel to theinner surface303 ofairfoil wall301 and/or theouter surface305 of theairfoil wall301. Additionally, it is contemplated that thesecond opening320 may be positioned in thepressure side wall306 or thesuction side wall308 of theairfoil300 for each of the correspondingpockets310,312,314, and316. Thesepockets310,312,314, and316 may be radially arrayed and fully enclosed within theairfoil wall301, having a pocket height in a radial direction to airfoil wall thickness at a minimum ratio of 1:1. Further, the positioning and structure ofpockets310,312,314, and316 may be manufactured using additive manufacturing.
• FIG. 4A is a cross-sectional view of anexemplary airfoil400, in accordance with an embodiment of the present invention. InFIG. 4A, theairfoil400 comprises anairfoil wall401, aleading edge402, aninner surface403, a trailingedge404, anouter surface405, apressure side wall406, and asuction side wall408. Theairfoil400 further includes a plurality ofairfoil passages410, which may allow cooling of theairfoil wall401 when cooling fluid or gas passes through theairfoil passages410.
• In theexemplary airfoil400, components of which are also shown inFIGS. 4B and 4C, theairfoil passages410 extend from theinner surface403 to theouter surface405 of theairfoil wall401 at various locations. Theairfoil passages410 in this embodiment allow cooling fluid or gas to enter arespective airfoil passage410 at afirst opening412, which may be one of a plurality offirst openings412, referred to hereinafter as thefirst opening412 for the sake of simplicity but intended to be non-limiting, and discharge the cooling fluid or gas from asecond opening414, which may be one of a plurality ofsecond openings414, referred to hereinafter as thesecond opening414 for the sake of simplicity but intended to be non-limiting. Achannel416 extends from thefirst opening412 to thesecond opening414 within theairfoil wall401.
• Additionally, inFIGS. 4A and 4B, a cross-sectional area of thechannel416 changes between thefirst opening412 and thesecond opening414. Theairfoil passage410 inFIGS. 4A-4C includes a cross-sectional area change between thefirst opening412 and thesecond opening414 that is approximately four to one; however, it is contemplated that the cross-sectional area difference may vary from 1.1:1 to 10:1 between the first and thesecond opening412,414, or have another relative difference. Theairfoil passage410 in thisairfoil400 is generally described as tapering in a radial direction, as the reduction in area between thefirst opening412 and thesecond opening414 occurs in the direction of cooling fluid flow along the radius of the rotor disk (not shown).
• FIG. 4C illustrates an enlarged perspective view of anairfoil passage410 having thefirst opening412 with a first cross-sectional area and thesecond opening414 with a second cross-sectional area that is smaller than the first-cross-sectional area. Additionally, thechannel416 further comprises afirst section418 having the first cross-sectional area along its axial length, asecond section420 having the second cross-sectional area along its axial length, and atransitional section422 having a cross-sectional area that tapers between the first cross-sectional area and the second cross-sectional area of the respective first andsecond sections418,420. Thetransitional section422 may taper linearly or non-linearly along the length of the transitional section422 (or any of the sections may taper). Thesecond section420 may further utilize a diffusion cooling hole to emit cooling fluid or gas from within theairfoil400 at high velocity and cause the emitted cooling fluid or gas to wrap over the outer surface of theairfoil400. This creates a thin, protective film layer of cooling fluid or gas between theouter surface405 of theairfoil400 and the high temperature combustion gases. A diffusion cooling hole may be utilized with theairfoil passage410 described herein, and the resulting outward cross-sectional area difference of thesecond section420 does not detract from the heat transfer coefficient benefits of a decreasing taper of thefirst section418 and thetransitional section422 of theairfoil passage410.
• Cooling fluid or gas entering thefirst section418 of theoperating airfoil400 may be relatively cool compared to theairfoil wall401. However, as cooling fluid or gas travels fromfirst section418 to thetransitional section422 and to thesecond section420, the cooling fluid or gas will gradually increase in temperature. Therefore, in order to provide a constant amount of heat transfer throughout the length of thechannel416, the cooling fluid or gas flow in thesecond section420 should travel at a higher velocity than the cooling fluid or gas flow through thefirst section418. As a result, the cross-sectional area ofsecond section420 is smaller than the cross-sectional area offirst section418 to increase the velocity of cooling fluid or gas traveling through theairfoil passage410.
• Additionally, as shown inFIG. 4C, afirst angle424 is formed between thefirst section418 and a correspondinginner surface403 of the airfoil wall401 (as shown inFIG. 4A), and may be between 15 and 90 degrees, and asecond angle426 is formed between thesecond section420 and theouter surface405 of the airfoil wall400 (as shown inFIG. 4A), which may be between 15 and 90 degrees. The taper of thetransitional section422 may generally occur in the radial direction of theairfoil wall401. However, thechannel416 may extend and/or taper in a radial and/or an axial direction ofairfoil wall401, or in another direction. Further, inFIG. 4C, thefirst section418, thesecond section420, and thetransitional section422 are shown generally in linear axial alignment. Alternatively,first section418,second section420, andtransitional section422 may be arranged in non-linearly.
• Thetransitional section422 may be oriented generally parallel to theairfoil wall401 and may be further characterized by a ratio of transitional section length to airfoil wall width. The airfoil wall width may be defined as the thickness between theinner surface403 of theairfoil wall401 and theouter surface405 of theairfoil wall401. The transitional section length, fully enclosed within an airfoil wall in a generally axial direction, to airfoil wall width may be a minimum ratio of 3:1 to a maximum ratio dependent upon an airfoil span between theleading edge402 and the trailingedge404 of theairfoil400.
• FIG. 4D. is a cross-sectional, perspective view of theairfoil passage410 incorporated into theairfoil400 shown inFIGS. 4A and 4B, in accordance with an embodiment of the present invention. InFIG. 4D,airfoil passage410 includes aflow turbulator428 within theairfoil passage410. The flow turbulator428 is shown as having a rectangular cross-section, but it is contemplated that theflow turbulator428 may have any uniform or non-uniform shape optimized for increasing the rate of convective heat transfer between theairfoil400 and the flow of cooling fluid or gas. Additionally, theflow turbulator428 may comprise a plurality offlow turbulators428 that may be arrayed in a linear or non-linear pattern within theairfoil passage410, or may be integrally manufactured with theairfoil passage410 to have a rough surface. In the disclosed embodiments, the heat transfer coefficient of theairfoil passage410 may be additionally modified by tailoring the surface roughness of the interior of theairfoil passage410 to achieve an equivalent roughness value of at least 400 Ra.
• FIG. 5A is an angled, cross-sectional, perspective view of anairfoil500 with variety ofairfoil passages510 integrated into anairfoil wall501 of theairfoil500, in accordance with an embodiment of the present invention. Theairfoil500 inFIG. 5A further comprises a leadingedge airfoil passage504 within theairfoil wall501, which extends at least partially onto the sides of theairfoil500.
• The leadingedge airfoil passage504 includes at least onefirst opening512 in theouter surface505 of theairfoil wall501, at least onesecond opening514 in theouter surface505 of theairfoil wall501, and achannel518 extending between thefirst opening512 and thesecond opening514 within theairfoil wall501. The leadingedge airfoil passage504 further comprises at least one third opening516 (which, inFIG. 5A, comprises two adjacent openings) in theinner surface503 of theairfoil wall501, which provides fluid communication between thechannel518 and anairfoil chamber507 at least partially enclosed by theairfoil wall501, through which cooling fluid or air may travel.
• The cross-sectional area of thechannel518 is largest adjacent or proximate thethird opening516 at a thirdcross-sectional area511 of thechannel518. Thethird opening516, which may supply cooling fluid or gas from theairfoil chamber507 to at least one of thefirst opening512 and thesecond opening514, and the thirdcross-sectional area511 of thechannel518, is positioned proximate a stagnation region of high temperature corresponding to leadingedge surface502. This positioning of thethird opening516 within thechannel518, betweenfirst opening512 andsecond opening514 near the thirdcross-sectional area511, allows the impingement effects of thethird opening516 to more effectively cool theairfoil wall501.
• The exemplary leadingedge airfoil passage504 may taper from the thirdcross-sectional area511 axially and/or radially towards thefirst opening512 and thesecond opening514 within theleading edge502 of theairfoil passage504 in order to accelerate the flow of cooling fluid or gas passing through the leadingedge airfoil passage504. The leadingedge airfoil passage504 may be duplicated across theleading edge502 of theairfoil500 to provide enhanced cooling across theleading edge502 of theairfoil500 during operation of the gas turbine.
• A first cross-sectional area of thefirst opening512, which may be one of a plurality offirst openings512, referred to hereinafter as thefirst opening512 for simplicity but intended to be non-limiting, of the leadingedge airfoil passage504 may be larger than a second cross-sectional area of thesecond opening514, which may be one of a plurality ofsecond openings514, referred to hereinafter as thesecond opening514 for simplicity but intended to be non-limiting, of the leadingedge airfoil passage504. The cross-sectional areas of thefirst opening512 andsecond opening514 are defined as the area between the walls of the channel at any position along the axial length of the channel. The leadingedge airfoil passage504 may be supplied with cooling fluid or gas from theairfoil chamber507 through thethird opening516 in theinner surface503 of theairfoil wall501. Thethird opening516, which may be one of a plurality ofthird openings516, referred to hereinafter as thethird opening516 for simplicity but intended to be non-limiting, may further be referred to as an impingement hole. This cooling fluid or gas enters theairfoil wall501 through thethird opening516, and then travels through thechannel518 towards thefirst opening512 and thesecond opening514 to exit the leadingedge airfoil passage504, carrying heat away from theairfoil wall501.
• The cross-sectional area of thechannel518 in the leadingedge airfoil passage504, as well as theother airfoil passages510, may vary, linearly or non-linearly, across the length ofchannel518, depending on the desired amount of heat transfer at different portions of the leadingedge airfoil passage504. In this respect, as shown in the leadingedge airfoil passage504, the cross-sectional area may be larger at the thirdcross-sectional area511 of thechannel518 than at the first andsecond openings512,514, to allow acceleration of cooling fluid or gas between thethird opening516 and the first andsecond openings512,514 during cooling of theairfoil500.
• FIG. 5B is a cross-sectional, elevation view of theairfoil500 ofFIG. 5A showing the plurality ofairfoil passages510 integrated therein, in accordance with an embodiment of the present invention. InFIG. 5B, as discussed with respect toFIG. 5A, the leadingedge airfoil passage504, which may be repeated along theleading edge502 of theairfoil500, may be supplied with cooling fluid or gas from theairfoil chamber507 through thethird opening516. This cooling fluid or gas travels through theleading edge502 of theairfoil500 by passing through thechannel518 tofirst opening512 andsecond opening514 to exit theairfoil wall501, carrying heat away from theairfoil500.
• FIG. 6 depicts a cut-out, perspective view of the geometry of a plurality of leadingedge airfoil passages604 integrated into anairfoil600, in accordance with an embodiment of the present invention.FIG. 6 is used to representatively show the three-dimensional geometry of the leadingedge airfoil passages604 as they are arrayed on theleading edge602 of theairfoil600. Furthermore, the leadingedge airfoil passages604 are connected via a plurality of connectingpassages609. The connectingpassages609 provide fluid communication between each of the plurality of leadingedge airfoil passages604. The connectingpassages609 may be positioned at any location along leadingedge602, in order to provide the desired fluid communication between each of the plurality of leadingedge airfoil passages604. Additionally, connectingpassages609 may be any shape, cross-sectional area, or frequency across the plurality of leadingedge airfoil passages604.
• FIGS. 7A-10C depict a variety ofairfoil passage geometries700,800,900,1000,1010, and1020 that can be integrated into an airfoil to provide enhanced cooling, in accordance with embodiments of the present invention. Referring now toFIGS. 7A-7C, a plurality ofchannels702 having generally sharp-edgedcorners704 are provided, in accordance with an embodiment of the present invention. The sharp-edgedcorners704 are generally formed when two ormore channels702 having different angles intersect. Additionally, the intersections ofchannels702 may be utilized to provide flow communication between thechannels702. Cooling fluid or gas may be supplied through thechannels702 via impingement holes706. The cooling fluid or gas may then exit thechannels702 throughopenings708 of therespective channels702. As previously discussed, thechannels702 may vary in cross-sectional area to control a velocity of cooling fluid or gas passing through thechannels702.
• FIG. 8 depicts a plurality ofchannels802 and803 in analternate arrangement800, in accordance with an embodiment of the present invention. InFIG. 8, cooling fluid or gas may be supplied to thechannels802 and803, with the channels separated by a dividingportion801. More specifically, the cooling fluid or air may be supplied to thechannels802 and803 through a plurality of impingement holes808, such that the cooling fluid or gas passes through thechannels802 and803 towards respective first andsecond openings810 and811. InFIG. 8, a plurality ofturbulators804 are shown along the length of a side-wall806 of thechannels802,803. The plurality ofturbulators804 are shown inFIG. 8 as having a rectangular cross-sectional shape. However, it is contemplated that the plurality ofturbulators804 may have other cross-sectional shapes, including asymmetrical or non-uniform shapes, or integrally manufactured leading edges having a rough surface. In the disclosed embodiments, the leading edge channel heat transfer coefficient may be modified by additionally tailoring the surface roughness to achieve an equivalent roughness of at least 400 Ra.
• As shown inFIG. 8, the plurality ofturbulators804 are arrayed in a parallel pattern along a length of thechannel802. However, the plurality ofturbulators804 may be patterned in a non-parallel pattern as well, in order to alter the fluid dynamics in thechannels802. For instance, theturbulators804 may comprise multiple rows of turbulators. Additionally, each row ofturbulators804 may be angled with respect to the channel802 (and anyother channels802,803 into which it is integrated). Further, turbulators804 may be positioned at any location withinchannels802 and803, and are not limited to a row configuration.
• Referring now toFIGS. 9A and 9B, a plurality of taperedchannels902 in analternate arrangement900 which may be integrated into a leading edge of an airfoil is provided, in accordance with an embodiment of the present invention. In operation, cooling fluid or gas may be provided to thechannels902 throughimpingement holes904 shown inFIGS. 9A and 9B. As cooling fluid or gas passes into thechannels902 from the impingement holes904, the cooling fluid or gas accelerates towards respectivefirst openings905 and respectivesecond openings907 of thechannels902 along aside wall906 due to the narrowing of thechannels902 towards theopenings905,907.
• Referring now toFIGS. 10A-10C,alternate arrangements1000,1010, and1020 of exemplary airfoil passages are depicted, in accordance with embodiments of the present invention. Thearrangements1000,1010, and1020 generally comprise different embodiments of a wave-like channel1002, which may be incorporated into a leading edge region of an airfoil. The wave-like channel1002, as shown inFIGS. 10A-10C, may comprise afirst portion1003 at a first angle, asecond portion1005 at a second angle, and a roundedtransitional portion1007 which connects the first and thesecond portions1003,1005. This roundedtransitional portion1007 creates the rounded “hill and valley” design effect shown inFIGS. 10A-10C. Such a pattern may be repeated throughout the wave-like channels1002. In operation, cooling fluid or gas may be provided to the plurality ofchannels1002 through impingement holes1004. As with prior designs, thechannels1002 may decrease in cross-sectional area from therespective impingement holes1004 to respective first andsecond openings1008,1009.
• Referring now toFIG. 11, a block diagram of anexemplary method1100 of manufacturing airfoils is provided, in accordance with an embodiment of the present invention. Atblock1110, an airfoil, such as theairfoil500 depicted inFIG. 5A, is provided. The airfoil comprises an airfoil wall, such as theairfoil wall501 shown inFIG. 5A, including an inner surface, such as theinner surface503 shown inFIG. 5A, and an outer surface, such as theouter surface505 shown inFIG. 5A, the airfoil wall forming an airfoil chamber, such as theairfoil chamber507 shown inFIG. 5A, at least partially enclosed within the airfoil wall.
• Atblock1120, a plurality of airfoil passages, such as the leadingedge airfoil passage504 shown inFIG. 5A, are formed at a leading edge, such as theleading edge502 of theairfoil500 shown inFIG. 5A, of the airfoil wall. As discussed herein, each of the plurality of airfoil passages comprises a first opening, such as thefirst opening512 shown inFIG. 5A, in the outer surface, a second opening, such as thesecond opening514 shown inFIG. 5A, in the outer surface, and a channel, such as thechannel518 shown inFIG. 5A, extending from at least one of the first opening and the second opening to a third opening, such as thethird opening516 shown inFIG. 5A, the third opening providing fluid communication between the channel and the airfoil chamber.
• The plurality of airfoil passages may be formed using additive manufacturing, such as selective laser melting (SLM), or another method. The first opening may include a first cross-sectional area and the second opening may include a second cross-sectional area, the first cross-sectional area being larger than the second cross sectional area.
• Referring now toFIG. 12, a block diagram of anotherexemplary method1200 of manufacturing airfoils is provided, in accordance with an embodiment of the present invention. Atblock1210, an airfoil, such as theairfoil500 depicted inFIG. 5A, is provided. The airfoil comprises an inner surface, such as theinner surface503 shown inFIG. 5A, and an outer surface, such as theouter surface505 shown inFIG. 5A, such that the airfoil wall forms an airfoil chamber, such as theairfoil chamber507 shown inFIG. 5A, at least partially enclosed within the airfoil wall. Atblock1220, a plurality of airfoil passages, such as theairfoil passages510 shown inFIG. 5A, are formed within the airfoil wall. Each of the airfoil passages comprises at least one first opening, such as thefirst opening512 shown inFIG. 5A, in the inner surface, at least one second opening, such as thesecond opening514 shown inFIG. 5A, in the outer surface, and a channel, such as thechannel518 shown inFIG. 5A, extending from the first opening to the second opening. The channel decreases in cross-sectional area between the at least one first opening and the at least one second opening. The plurality of airfoil passages may be formed at least partially in a leading edge wall of the airfoil, and/or at least partially on a pressure side wall and a suction side wall of the airfoil.
• Referring now toFIG. 13, a block diagram of anotherexemplary method1300 of manufacturing airfoils is provided, in accordance with an embodiment of the present invention. Atblock1310, an airfoil, such as theairfoil500 shown inFIG. 5A, having a leading edge, such as theleading edge502 shown inFIG. 5A, and a trailing edge, such as the trailingedge404 shown inFIG. 4A, is provided. The airfoil comprises an airfoil wall, such as theairfoil wall501 shown inFIG. 5A, having an inner surface, such as theinner surface503 shown inFIG. 5A, and an outer surface, such as theouter surface505 shown inFIG. 5A, the airfoil wall forming an airfoil chamber, such as theairfoil chamber507 shown inFIG. 5A, at least partially enclosed by the airfoil wall.
• Atblock1320, a plurality of pockets, such as thepockets310,312,314, and316 shown inFIG. 3A, are formed within the airfoil wall. Each of the plurality of pockets comprises an inner pocket wall, such as theinner pocket wall324 shown inFIG. 3A, and an outer pocket wall, such as theouter pocket wall326 shown inFIG. 3A. Additionally, a first opening, such as thefirst opening318 shown inFIG. 3A, may be positioned in the inner surface at a first distance away from the leading edge, the first opening providing fluid communication between the airfoil chamber and the pocket, and a second opening, such as thesecond opening320 shown inFIG. 3A, may be positioned at a second distance away from the leading edge, the second opening providing fluid communication between an outside of the airfoil and the pocket. Further, a distance between the inner pocket wall and the outer pocket wall is larger proximate the leading edge of the airfoil and smaller approximate the trailing edge of the airfoil.
• From the foregoing, it will be seen that this invention is one well adapted to attain all the ends and objects hereinabove set forth together with other advantages which are obvious and which are inherent to the structure. It will be understood that certain features and subcombinations are of utility and may be employed without reference to other features and subcombinations. This is contemplated by and is within the scope of the claims. Since many possible embodiments may be made of the invention without departing from the scope thereof, it is to be understood that all matter herein set forth or shown in the accompanying drawings is to be interpreted as illustrative and not in a limiting sense. Additional objects, advantages, and novel features of the invention will be set forth in part in the description which follows, and in part will become apparent to those skilled in the art upon examination of the following, or may be learned by practice of the invention.

Claims (25)

What is claimed is:

1. An airfoil for a gas turbine having a leading edge and a trailing edge, the airfoil comprising:

an airfoil wall having an inner surface and an outer surface, the airfoil wall forming an airfoil chamber at least partially enclosed by the airfoil wall; and

a plurality of pockets within the airfoil wall, each of the plurality of pockets comprising:

an inner pocket wall and an outer pocket wall;

at least one first opening in the inner surface positioned at a first distance away from the leading edge, the first opening providing fluid communication between the airfoil chamber and the pocket; and

at least one second opening in the outer surface positioned at a second distance away from the leading edge, the second opening providing fluid communication between an outside of the airfoil and the pocket,

wherein a distance between the inner pocket wall and the outer pocket wall is larger proximate the leading edge and smaller proximate the trailing edge.

2. The airfoil ofclaim 1, wherein each of the plurality of pockets further comprises a plurality of pedestals, wherein the plurality of pedestals extend between the inner pocket wall and outer pocket wall of each respective pocket of the plurality of pockets.

3. The airfoil ofclaim 2, wherein each pedestal of the plurality of pedestals of each of the plurality of pockets has a circular, triangular, square, ovular, or rectangular cross-section.

4. The airfoil ofclaim 2, wherein a distance between each of the plurality of pedestals in each of the plurality of pockets is larger proximate the leading edge and smaller proximate the trailing edge.

5. The airfoil ofclaim 2, wherein the plurality of pockets are arrayed in a linear or non-linear pattern within the airfoil wall.

6. The airfoil ofclaim 2, wherein the plurality of pockets within the airfoil wall and the plurality of pedestals are formed using additive manufacturing.

7. The airfoil ofclaim 1, wherein a ratio of cross-sectional area between the inner pocket wall and the outer pocket wall of each of the plurality of pockets is between 1.1:1 and 10:1.

8. The airfoil ofclaim 7, wherein for each of the plurality of pockets, the inner pocket wall and the outer pocket wall are substantially parallel to the inner surface of the airfoil wall.

9. The airfoil ofclaim 1, wherein for each of the plurality of pockets, the second opening is positioned on a pressure side wall or a suction side wall of the airfoil.

10. The airfoil ofclaim 1, wherein a ratio of pocket axial length to airfoil wall width in each of the plurality of pockets is at least 3:1.

11. The airfoil ofclaim 1, wherein the plurality of pockets are arrayed in a radial direction within the airfoil wall, and wherein a ratio of pocket radial height to airfoil wall width in each of the plurality of pockets is at least 3:1.

12. A gas turbine assembly, the assembly comprising:

an airfoil having a leading edge and a trailing edge, the airfoil comprising an airfoil wall having an inner surface and an outer surface, the airfoil wall forming an airfoil chamber at least partially enclosed by the airfoil wall; and

a plurality of pockets within the airfoil wall, each of the plurality of pockets comprising:

an inner pocket wall and an outer pocket wall;

at least one first opening in the inner surface positioned at a first distance away from the leading edge, the first opening providing fluid communication between the airfoil chamber and the pocket; and

at least one second opening in the outer surface positioned at a second distance away from the leading edge, the second opening providing fluid communication between an outside of the airfoil and the respective pocket,

wherein a distance between the inner pocket wall and the outer pocket wall is larger proximate the leading edge and smaller proximate the trailing edge, and

wherein a ratio of cross-sectional area of the inner pocket wall to the outer pocket wall is between 1.1:1 and 10:1.

13. The assembly ofclaim 12, wherein each of the plurality of pockets further comprises a plurality of pedestals that extend between the inner pocket wall and the outer pocket wall.

14. The assembly ofclaim 13, wherein each pedestal of the plurality of pedestals of each of the plurality of pockets has a circular, triangular, square, ovular, or rectangular cross-section.

15. The assembly ofclaim 14, wherein a distance between each of the plurality of pedestals in each of the plurality of pockets is larger proximate the leading edge and smaller proximate the trailing edge.

16. The assembly ofclaim 12, wherein the plurality of pockets within the airfoil wall are arrayed in a linear or a non-linear pattern.

17. The assembly ofclaim 12, wherein the inner pocket wall and the outer pocket wall of each of the plurality of pockets are substantially parallel to the inner surface of the airfoil wall.

18. The assembly ofclaim 12, wherein the second opening is positioned in a pressure side wall, a suction side wall, or the trailing edge of the airfoil.

19. The assembly ofclaim 12, wherein a ratio of pocket axial length to airfoil wall width in each of the plurality of pockets is at least 3:1.

20. The assembly ofclaim 12, wherein a ratio of pocket radial height to airfoil wall width in each of the plurality of pockets is at least 3:1.

21. A method of manufacturing airfoils, the method comprising:

providing an airfoil having a leading edge and a trailing edge, the airfoil having an airfoil wall having an inner surface and an outer surface, the airfoil wall forming an airfoil chamber at least partially enclosed by the airfoil wall; and

forming a plurality of pockets within the airfoil wall, each of the plurality of pockets comprising:

an inner pocket wall and an outer pocket wall;

a first opening in the inner surface positioned at a first distance away from the leading edge, the first opening providing fluid communication between the airfoil chamber and the pocket; and

a second opening in the outer surface positioned at a second distance away from the leading edge, the second opening providing fluid communication between an outside of the airfoil and the pocket,

wherein a distance between the inner pocket wall and the outer pocket wall is larger proximate a leading edge of the airfoil and smaller proximate a trailing edge of the airfoil.

22. The method ofclaim 21, wherein the pocket is formed using additive manufacturing.

23. The method ofclaim 21, wherein a ratio of pocket axial length to airfoil wall width for each of the plurality of pockets is at least 3:1.

24. The method ofclaim 21, wherein a ratio of pocket radial height to airfoil wall width for each of the plurality of pockets is at least 3:1.

25. The method ofclaim 21, wherein each of the plurality of pockets further comprises a plurality of pedestals, wherein the plurality of pedestals extend between the inner pocket wall and outer pocket wall of each respective pocket of the plurality of pockets.

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