Two-part cooling channel for airfoils

Abstract

This invention relates to a turbine airfoil comprising a body having airfoil walls and a coolant chamber within the body. A plurality of cooling channels are located within the airfoil walls, each cooling channel comprising: a first portion extending from a first point on an outer surface of the airfoil wall to the coolant chamber; and a second portion extending from a second point on the outer surface of the airfoil wall, away from the first point, to intersect with a middle portion of the first portion. A cover element closes the first portion at the first point, but leaves the second portion open. Each cooling channel has a single inlet in fluid communication with the coolant chamber and a single outlet at the second point of the second portion.

Description

[0001] Cross-reference to related applications This application relates to concurrently filed and currently pending U.S. application US 15 / 862927 (GE Case No. 323936-1). Technical Field

[0002] This disclosure generally relates to turbine airfoil cooling, and more specifically to a two-part cooling channel for turbine airfoils. Background Technology

[0003] Gas turbine blades' airfoils and nozzles are exposed to excessive thermal loads. Airfoils are typically covered with a high concentration of thermal barrier coating (TBC). Therefore, the TBC undergoes delamination (or spalling), making the airfoil more difficult to cool. To cool the airfoil, coolant is typically introduced from the airfoil's inner chamber through orifices into the airfoil's outer surface via cooling channels. A large number of cooling channels are arranged, creating numerous orifices within the airfoil. At the leading edge, this arrangement of cooling orifices can be referred to as a showerhead arrangement. Ideally, the coolant generates a cooling film (i.e., a flow across and close to the airfoil's surface) that extends downstream along the airfoil's surface.

[0004] On the leading edge, cooling channels with conventional circular or conical exit holes are radially oriented relative to the surface; that is, these cooling channels are drilled perpendicular to the direction of hot air flow. Therefore, the cooling flow must make a sharp turn and is susceptible to blow-off from the airfoil surface, which can reduce coolant coverage and lateral average cooling effectiveness. Cooling channels with shaped diffuser exit holes are typically used in other areas of the airfoil and have relatively high cooling effectiveness, but have not been successfully used in the leading edge due to its small radius of curvature. That is, shaped diffuser exit holes need to be drilled almost perpendicular to the surface using conventional manufacturing methods. This arrangement results in reduced cooling performance.

[0005] In addition to the challenges mentioned above, there is an increasing need to reduce overall coolant usage while maintaining membrane cooling at the leading edge to achieve forward tilt efficiency. To reduce overall cooling flow, fewer cooling channels are typically used, resulting in increased spacing between the channels. Each cooling channel must therefore have improved lateral average cooling effectiveness and increased lateral coolant coverage. Many advanced membrane designs rely on additive manufacturing to generate improved cooling channels, but several challenges need to be addressed before this technology can be widely deployed in hot gas path segments. Summary of the Invention

[0006] A first aspect of this disclosure provides a turbine airfoil comprising: a body having an airfoil wall; a coolant chamber within the body; and a plurality of cooling channels within the airfoil wall, each cooling channel comprising: a first portion extending from a first point on an outer surface of the airfoil wall to the coolant chamber; a second portion extending from a second point on the outer surface of the airfoil wall away from the first point to intersect with a middle portion of the first portion; and a cover element that closes the first portion at the first point but leaves the second portion open, wherein each cooling channel has a single inlet in fluid communication with the coolant chamber and a single outlet at the second point of the second portion.

[0007] A second aspect of this disclosure provides a method for forming cooling channels in the airfoil wall of a turbine airfoil, the airfoil including a coolant chamber therein, the method comprising: forming a first portion of the cooling channel extending from a first point on an outer surface of the airfoil wall to the coolant chamber; forming a second portion of the cooling channel extending from a second point on the outer surface of the airfoil wall away from the first point to intersect with a middle portion of the first portion; and closing the first portion at the first point using a cap element, but leaving the second portion open, wherein each cooling channel has a single inlet in fluid communication with the coolant chamber and a single outlet at the second point of the second portion.

[0008] The illustrative aspects of this disclosure are designed to address the problems described herein and / or other problems not discussed herein.

[0009] Technical Solution 1. A turbine airfoil component, comprising: A main body with airfoil-shaped walls; The coolant chamber within the main body; and The airfoil wall contains multiple cooling channels, each cooling channel including: The first part extends from a first point on the outer surface of the airfoil wall to the coolant chamber. The second part extends from a second point on the outer surface of the airfoil wall, away from the first point, to intersect with the middle portion of the first part, and A cover element that closes the first portion at the first point, but leaves the second portion open. Each cooling channel has a single inlet in fluid communication with the coolant chamber and a single outlet at a second point in the second section.

[0010] Technical Solution 2. The turbine airfoil according to Technical Solution 1, wherein the second point of the second part includes a diffuser opening in the axially downstream direction from the leading edge.

[0011] Technical Solution 3. The turbine airfoil according to Technical Solution 2, wherein the second portion diverges relative to the outer and inner surfaces of the airfoil wall along its length from the middle portion of the first portion to the second point.

[0012] Technical Solution 4. The turbine airfoil according to Technical Solution 2, wherein the second portion radially diverges relative to the length of the body at an angle between 0° and approximately 20° along the length of the second portion from the middle portion of the first portion to the second point.

[0013] Technical Solution 5. The turbine airfoil according to Technical Solution 1, wherein the second part intersects with the middle part of the first part.

[0014] Technical Solution 6. The turbine airfoil according to Technical Solution 1, wherein the first portion meets the outer surface of the airfoil wall at a first point at an angle between approximately 45° and approximately 90°, and the second portion meets the outer surface of the airfoil wall at a second point at an angle between approximately 135° and approximately 175°.

[0015] Technical Solution 7. The turbine airfoil according to Technical Solution 1, wherein the cover element includes a thermal barrier coating (TBC), and the second portion of each cooling channel extends through the TBC at its second point.

[0016] Technical Solution 8. The turbine airfoil according to Technical Solution 1, wherein the cover element includes a metal plug at or near the first point in the first portion, and, It also includes a thermal barrier coating (TBC) covering the outer surface of the airfoil wall, with the second portion of each cooling channel extending through the TBC at its second point.

[0017] Technical Solution 9. The turbine airfoil according to Technical Solution 1, wherein the cover element includes a metal layer that traverses the first portion at the first point, and, It also includes a thermal barrier coating (TBC) covering the outer surface of the airfoil wall, with the second portion of each cooling channel extending through the TBC at its second point.

[0018] Technical Solution 10. The turbine airfoil according to Technical Solution 1, wherein a portion of the airfoil includes the plurality of cooling channels, and the cover element includes a hanger attached to the remainder of the body.

[0019] Technical Solution 11. The turbine airfoil according to Technical Solution 1, wherein the cross-sectional area of ​​the first part is smaller than the cross-sectional area of ​​the second part, and the second part intersects with the middle part of the first part.

[0020] Technical Solution 12. The turbine airfoil according to Technical Solution 1, wherein the body includes a base and a tip arranged relative to the base, and a pressure sidewall and a suction sidewall extend between a leading edge and a trailing edge, wherein the airfoil wall is arranged in the leading edge.

[0021] Technical Solution 13. A method for forming a cooling channel in the airfoil wall of a turbine airfoil, the turbine airfoil including a coolant chamber therein, the method comprising: A first portion forming the cooling channel extends from a first point on the outer surface of the airfoil wall to the coolant chamber; A second portion forming the cooling channel extends from a second point on the outer surface of the airfoil wall, away from the first point, to intersect with the middle portion of the first portion; and At the first point, a cover element is used to close the first part, but the second part remains open. Each cooling channel has a single inlet in fluid communication with the coolant chamber and a single outlet at the second point of the second section.

[0022] Technical Solution 14. The method according to Technical Solution 13, wherein each formation includes drilling.

[0023] Technical Solution 15. The method according to Technical Solution 13, wherein forming the first portion includes drilling outward from the coolant chamber.

[0024] Technical Solution 16. The method according to Technical Solution 13, wherein forming the second portion includes forming a second point of the second portion to include a diffuser opening in an axially downstream direction from the leading edge.

[0025] Technical Solution 17. The method according to Technical Solution 13, wherein using the cover element to close the first portion includes forming a thermal barrier coating (TBC) throughout the outer surface, and the second portion of each cooling channel extends through the TBC at its second point.

[0026] Technical Solution 18. The method according to Technical Solution 13, wherein using the cover element to close the first portion includes a metal plug forming the first point in or near the first portion, and, It also includes forming a thermal barrier coating (TBC) all over the outer surface, with the second portion of each cooling channel extending through the TBC at its second point.

[0027] Technical Solution 19. The method according to Technical Solution 13, wherein closing the first portion using the cover element includes a metal layer traversing the first portion at the first point, and, It also includes forming a thermal barrier coating (TBC) all over the outer surface, with the second portion of each cooling channel extending through the TBC at its second point.

[0028] Technical Solution 20. The method according to Technical Solution 13, wherein forming the first portion and the second portion includes forming a cross-sectional area of ​​the first portion that is smaller than the cross-sectional area of ​​the second portion, wherein the second portion intersects with the middle portion of the first portion. Attached Figure Description

[0029] These and other features of the present disclosure will be more readily understood from the following detailed description of various aspects of the present disclosure, which are jointly depicted in the accompanying drawings illustrating various embodiments thereof, wherein: Figure 1 A schematic diagram of a turbomachinery system in the form of a gas turbine system is shown.

[0030] Figure 2 It shows that it can be used with Figure 1 An illustrative cross-sectional view of a gas turbine assembly used in a gas turbine system.

[0031] Figure 3 A perspective view of a turbine rotor blade of a type that may employ embodiments of this disclosure is shown.

[0032] Figure 4 A perspective view of a turbine guide vane of a type that may employ embodiments of this disclosure is shown.

[0033] Figure 5 A partial perspective view of a turbine airfoil according to an embodiment of the present disclosure is shown.

[0034] Figure 6 A partial cross-sectional view of the airfoil wall of a turbine airfoil according to an embodiment of the present disclosure is shown.

[0035] Figure 7 Show Figure 6 A magnified partial cross-sectional view of the leading edge.

[0036] Figure 8 An enlarged partial cross-sectional view of the airfoil wall of a turbine airfoil according to another embodiment of the present disclosure is shown.

[0037] Figure 9 Embodiments according to this disclosure are shown Figure 7 Or an enlarged planar view of the leading edge of 8.

[0038] Figure 10 An enlarged partial cross-sectional view of the airfoil wall of a turbine airfoil according to another embodiment of the present disclosure is shown.

[0039] Figure 11 An enlarged partial cross-sectional view of the airfoil wall of a turbine airfoil according to another embodiment of the present disclosure is shown.

[0040] Figure 12 An enlarged partial cross-sectional view of the airfoil wall of a turbine airfoil according to another embodiment of the present disclosure is shown.

[0041] Figure 13 An enlarged partial cross-sectional view of the airfoil wall of a turbine airfoil according to another embodiment of the present disclosure is shown.

[0042] It should be noted that the accompanying drawings of this disclosure are not drawn to scale. The drawings are intended to depict only typical aspects of this disclosure and therefore should not be considered as limiting the scope of this disclosure. In the drawings, the same numbers denote the same elements between the figures.

[0043] Reference Symbol List Turbomachinery 100 Compressor 102 Burner 104 Combustion Zone 106 Fuel nozzle assembly 108 Turbine assembly 110 Turbo 111 Rotor 112 Guide vane 120 Fixed outer casing 122 Rotary blade 124 Guide vane 126 External platform 128 Internal platform 130 Turbine rotor blades 132 Inward Platform 134 Outward-facing pointed shield 136 Root 140 Dovetail part 142 Rotor wheel 144 Handle 146 Platform 148 Airfoil 150 Concave pressure sidewall 152 Relative convex suction sidewall 154 Front edge 156 posterior edge 158 160 degrees on the outer tip Fixed guide vane 170 External platform 172 Internal platform 174 Airfoil 176 Concave pressure sidewall 178 The opposite convex suction sidewalls 180 Frontier 182 posterior edge 184 200 airfoil Cooling channel 204 Main body 210 Frontier 212 Airfoil wall 214 Coolant chamber 216 Part 1, Chapter 218 Point 1, 220 outer surface 222 Part Two, 230 Second point 232 Middle section 234 Diffuser opening 240 Inner surface 242 Cover element 250 252 single entrances Exports 254 Thermal barrier coating TBC 260 Adhesive coating layer 262 TBC layer 264 Metal plug 270 280 brazed metal layers Relief cut 282 286 brazed metal parts Enlarged entrance section 290 Hanging piece 292. Detailed Implementation

[0044] First, when referring to and describing relevant machine components within a gas turbine system, it will be necessary to choose a term in order to clearly describe this disclosure. When choosing a term, it will be used and adopted in a manner consistent with the generally accepted meaning of common industrial terms, where possible. Unless otherwise stated, such terms should be given a broad interpretation consistent with the context of this application and the scope of the appended claims. Those skilled in the art will recognize that, generally, several different or overlapping terms may be used to refer to a particular component. A portion that can be described herein as a single part may include multiple components, and, in another context, is referred to as consisting of multiple components. Alternatively, a portion that can be described herein as comprising multiple components may elsewhere be referred to as a single part.

[0045] In addition, several descriptive terms may be used conventionally herein, and should prove helpful in defining these terms at the beginning of this paragraph. Unless otherwise stated, these terms and their definitions are as follows. As used herein, “downstream” and “upstream” are terms indicating the direction of flow relative to a fluid (such as the working fluid through a turbine engine, or, for example, the airflow through a combustor or the coolant through one of the components of a turbine). The term “downstream” corresponds to the direction of fluid flow, and the term “upstream” refers to the direction opposite to that flow. Without any further specificity, the terms “front” and “rear” refer to directions, where “front” refers to the front of the engine or the compressor end, and “rear” refers to the rear of the engine or the turbine end. It is often required to describe parts with respect to different radial positions of the central axis. The term “radial” refers to movement or position perpendicular to the axis. In such cases, if the first component is located closer to the axis than the second component, it will be stated herein that the first component is located “radially inward” or “inner” than the second component. On the other hand, if the first component is located further away from the axis than the second component, then the position of the first component relative to the second component as "radially outward" or "outer" can be specified in this document. The term "axial" refers to movement or position parallel to the axis. Finally, the term "circumferential" refers to movement or position about the axis. It will be appreciated that such terms can be applied with respect to the central axis of the turbine.

[0046] Embodiments of this disclosure provide a turbine airfoil including cooling channels having: a first portion extending from a first point on the outer surface of the airfoil wall to a coolant chamber within the turbine airfoil, and a second portion extending from the second point on the outer surface of the airfoil wall at a second point away from the first point to intersect with a middle portion of the first portion. A cap element closes the first portion at the first point but leaves the second portion open. Thus, each cooling channel has a single inlet in fluid communication with the coolant chamber and a single outlet at the second point of the second portion. The first and second portions can be fabricated in two separate manufacturing steps. The outlet may include a diffuser to help guide the coolant. The turbine airfoil exhibits improved average film cooling effectiveness and increased coolant coverage compared to conventional showerhead arrangements, which can improve combined cycle efficiency and improve part durability. In other locations, the cooling channels can also be applied to the leading edge wall of the turbine airfoil.

[0047] Figure 1 This illustration shows a schematic diagram of an industrial machine to which the teachings of this disclosure can be applied. In this example, the machine includes a turbomachinery 100 in the form of a combustion or gas turbine system. Turbomachinery 100 includes a compressor 102 and a combustor 104. Combustor 104 includes a combustion zone 106 and a fuel nozzle assembly 108. Turbomachinery 100 also includes a turbine assembly 110 and a common compressor / turbine shaft (sometimes referred to as rotor 112). In one embodiment, the combustion turbine system is the MS7001FB engine (sometimes referred to as the 7FB engine) commercially available from General Electric Company in Greenville, South Carolina. This disclosure is neither limited to any particular industrial machine nor to any particular gas turbine system and can be incorporated in conjunction with other engines, including, for example, General Electric's MS7001FA (7FA), MS9001FA (9FA), 7HA, and 9HA engine models. Furthermore, this disclosure is not limited to any specific turbomachinery, but can be applied to any turbine airfoil that requires film cooling of the leading edge of the turbine airfoil, such as in steam turbines, jet engines, compressors, turbofan engines, etc.

[0048] During operation, air flows through compressor 102, and compressed air is supplied to burner 104. Specifically, compressed air is supplied to fuel nozzle assembly 108, which is integral with burner 104. Assembly 108 is in flow communication with combustion zone 106. Fuel nozzle assembly 108 is also in communication with fuel source (not in combustion zone 106). Figure 1(As shown in the diagram) A flow communication is established, directing fuel and air to combustion zone 106. Combustor 104 ignites and combusts the fuel. Combustor 104 is in flow communication with turbine assembly 110, whereby the thermal energy of the airflow is converted into mechanical rotational energy. Turbine assembly 110 includes a turbine rotatably coupled to and drives rotor 112. Compressor 102 is similarly rotatably coupled to rotor 112. In the illustrated embodiment, multiple combustion zones 106 and fuel nozzle assembly 108 are present.

[0049] Figure 2 It shows that it can be used with Figure 1 Turbomachinery 100 used in gas turbine systems Figure 1 The diagram shows a cross-sectional view of a turbine assembly 110. The turbine 111 of the turbine assembly 110 includes a row of nozzles or guide vanes 120 coupled to a fixed housing 122 of the turbine machinery 100 and axially adjacent to a row of rotating blades 124. The nozzles or guide vanes 126 may be held within the turbine assembly 110 by a radially outer platform 128 and a radially inner platform 130. The row of rotating blades 124 in the turbine assembly 110 includes turbine rotor blades 132 coupled to and rotating with a rotor 112. The turbine rotor blades 132 may include a radially inward platform 134 (at the root of the blade) (coupled to the rotor 112) and a radially outward tip guard 136 (at the tip of the blade).

[0050] Figure 3 and Figure 4 The illustration shows a hot gas path turbine component (including turbine airfoil) of a turbine machine that can adopt the teachings of this disclosure. Figure 3 A perspective view of a turbine rotor blade 132 of a type that may employ embodiments of the present disclosure is shown. The turbine rotor blade 132 includes an airfoil 150 having a body 151 with a root or base 140, the turbine rotor blade 132 being attached to a rotor 112 via the root or base 140. Figure 2 The base 140 may include a dovetail member 142, which is configured for mounting on the rotor 112. Figure 2 ) rotor wheel 144 ( Figure 2The base 140 may further include a shank 146 extending between the dovetail member 142 and the platform 148, the shank 146 being disposed at the junction of the airfoil 150 and the base 140 and defining a portion of the inner boundary of the flow path through the turbine assembly 110. It will be appreciated that the airfoil 150 is a moving member of the turbine rotor blades 132 that intercepts the flow of working fluid and causes the rotor disk to rotate. It will be seen that the body 151 of the airfoil 150 includes a concave pressure sidewall (PS) 152 and circumferentially or laterally opposed convex suction sidewalls (SS) 154, which extend axially between opposing leading edges 156 and trailing edges 158, respectively. The concave pressure sidewall 152 and the convex suction sidewall 154 similarly extend radially from the platform 148 to the outer tip 160.

[0051] Figure 4 A perspective view of a fixed guide vane 170 of a type that may employ embodiments of the present disclosure is shown. The fixed guide vane 170 includes an outer platform 172, through which the fixed guide vane 170 is attached to a fixed housing 122 of the turbomachinery. Figure 2 The outer platform 172 may include any currently known or later developed mounting configuration for mounting to a corresponding mounting element within the housing. The fixed guide vane 170 may further include an inner platform 174 (similar to the tip 160). Figure 3 In order to locate platform 148 ( Figure 3 The adjacent turbine rotor blade 132 () Figure 3 Platforms 172 and 174 define corresponding portions of the outer and inner boundaries of the flow path through turbine assembly 110. It will be appreciated that airfoil 176 is a fixed guide vane 170 that intercepts the flow of the working fluid and directs the working fluid toward turbine rotor blades 132. Figure 3The airfoil 176 of the fixed guide vane 170 includes a body 171 having a concave pressure sidewall (PS) 178 and circumferentially or laterally opposed convex suction sidewalls (SS) 180, which extend axially between opposing leading edges 182 and trailing edges 184, respectively. The concave pressure sidewall 178 and convex suction sidewalls 180 also extend radially from platform 172 to platform 174. Embodiments of this disclosure described herein may include aspects of turbine airfoils suitable for turbine rotor blades 132 and / or fixed guide vanes 170. It is understood that other features of the turbine rotor blades 132 or guide vanes 170 not described herein (such as, but not limited to, internal cooling structures, cutout shapes, outer wall angle configurations / shapes, etc.) may be tailored to the specific application (i.e., rotor blades or guide vanes).

[0052] refer to Figure 5-13 Embodiments of the turbine airfoil 200 according to embodiments of the present disclosure will now be described. As will be appreciated, the structure described with respect to airfoil 200 can be employed in conjunction with airfoils 150, 176 as previously described herein, as applied to gas turbine systems using turbine airfoils requiring cooling or any other industrial machinery. Figure 5 A partial perspective view is shown, and Figure 6 A partial cross-sectional view of the leading edge region of a turbine airfoil 200 employing a cooling channel 204 according to an embodiment of the present disclosure is shown. It will be understood that the turbine airfoil 200 may include (where necessary, refer to...) Figure 3 For reference to airfoil 150), the body 210 has a base 140 and a tip 160 arranged opposite the base. The airfoil wall 214 of the body 210 may include a concave pressure sidewall 152 and a convex suction sidewall 154 extending between a leading edge 156 and a trailing edge 158. According to some embodiments, the airfoil wall 214 (in which cooling channels are applied according to embodiments of the present disclosure) is arranged at the leading edge 156 ( Figure 3(as leading edge wall). As noted, the teachings of this disclosure can be applied to other locations of the airfoil wall 214 in the body 210. In some embodiments, the body 210 may comprise metal or metal alloys depending on the application of the turbine airfoil. Some examples include, but are not limited to: Rene 108, CM247, Haynes alloy, Incalloy, MP98T, TMS alloy, CMSX single crystal alloy. In other embodiments, the body 210 may comprise ceramic matrix composite (CMC). As understood, a coolant chamber 216 may pass through the body 210. The coolant chamber 216 may carry any form of coolant (e.g., air from any source (e.g., compressor or other coolant chamber)). As understood in the art, the coolant chamber 216 can take a wide variety of forms and shapes within the body 210. The body 210 and the coolant chamber 216 may be formed using any currently known or later developed processes for forming the turbine airfoil body (such as, but not limited to, casting and additive manufacturing).

[0053] like Figure 5 As shown, the turbine airfoil 200 may also include a plurality of cooling channels 204 within the airfoil wall 214. Embodiments of this disclosure are advantageous for the leading edge wall of the airfoil 200, which not only exhibits the characteristics of… Figure 5 and Figure 6 The small radius shown is applicable, and in fact, to any location within the airfoil wall 214. In one example, the leading edge may have a radius of curvature to wall thickness ratio of less than eight (8); other definitions of the portion constituting the leading edge may similarly be applicable. Furthermore, embodiments may be particularly advantageous near the radially inward and radially outward ends of the airfoil leading edge (i.e., the airfoil base and tip), where conventional radial holes require angles to the edges, creating areas of lower cooling efficiency. Figure 6 and Figure 7 As shown in the enlarged partial cross-sectional view, each cooling channel 204 may include a first portion 218 extending from a first point 220 on the outer surface 222 of the airfoil wall 214 to the coolant chamber 216. The first portion 218 may be formed using any currently known or later developed method of forming a hole through the airfoil wall 214. In one embodiment, as will be described herein, the first portion 218 is formed by drilling before the thermal barrier coating (TBC) 260 (e.g., an adhesive coating 262 with a TBC layer 264 or a TBC layer 264 alone) is formed on the outer surface 222. Drilling may occur (or be performed, i.e., occur) from the outside of the airfoil wall 214 inward to the coolant chamber 216. Alternatively, drilling may occur from the inside of the coolant chamber 216 outward (i.e., through the outer surface 222). Figure 7As illustrated, the first portion 218 can meet the outer surface 222 of the airfoil wall 214 at a first point 220 at an angle α between approximately 45° and approximately 90°.

[0054] Each cooling channel 204 may further include a second portion 230 extending from a second point 232 on the outer surface 222 of the airfoil wall 214, away from the first point 220, to intersect with a middle portion 234 of the first portion 218. The second point 232 may be located downstream of the first point 220 (i.e., the first portion 218 and the second portion 230 are separated at the outer surface 222). The middle portion 234 may be located anywhere between the ends of the first portion 218, i.e., such that the second portion 230 is not fluidly connected at its inner end (other than to the first portion 218). The second portion 230 may be formed using any currently known or later developed method of forming a hole through the airfoil wall 214. In one embodiment, as will be described herein, the second portion 230 may be formed by drilling before the TBC 260 is formed on the outer surface 222. Alternatively, the second portion 230 can be formed after the TBC 260 is formed on the outer surface 222, for example, by drilling through the TBC 260. In any case, the drilling can occur from the outside of the airfoil wall 214 inwards to intersect with the middle portion 234 of the first portion 218. The depth of the second portion 230 can be such that... Figure 6 and Figure 7 As shown, the second part 230 intersects with the first part 218, or as for example... Figure 8 As shown, the second part 230 can be simply connected to the first part 218. (As illustrated...) Figure 7 As illustrated, the second portion 230 can meet the outer surface 222 of the airfoil wall 214 at a second point 232 at an angle β between approximately 135° and approximately 175°. Compared to conventional showerhead holes, the angle β and the orientation of the second portion 230 generate a larger coolant coverage and improved average effectiveness. The cooling channel 204 can guide coolant to the pressure side and / or suction side of the turbine airfoil 200. Figure 8 and Figure 9As shown, the second portion 230 may optionally include a diffuser opening 240 at the second point 232 in an axially downstream direction from the leading edge 212. That is, the second portion 230 includes, at the second point 232 and perhaps preceding (or “leading up to”) the second point 232, the formation of a diffuser opening 240 in an axially downstream direction from the leading edge 212. The diffuser opening 240 includes a gradually widening wall that generates a gradually increasing cross-sectional area. Here, the second portion 230 diverges along its length from the middle portion 234 of the first portion 218 to the second point 232 relative to the outer surface 222 and inner surface 242 of the airfoil wall 214. Figure 9 As shown, each diffuser opening 240 may also have a second portion 230 such that the length of the second portion 230 from the middle portion 234 of the first portion 218 to the second point 232 is radially diverging relative to the length of the body at an angle between 0° and 20°. That is, each side of the diffuser opening 240 may form a radial angle δ of up to + / -0° to approximately 20°. It is also noted that when the diffuser opening 240 is provided, its cross-sectional area can be much larger than the cross-sectional area of ​​the first portion 218. Alternatively, the diffuser opening 240 may be omitted, and the second portion 230 may simply be formed as a hole, which will appear similar in cross-section to Figure 7 The cross-section.

[0055] In one embodiment, such as Figure 9 As shown in the plan view of the leading edge 212, the second portion 230 (illustrated within the outer surface 222) can radially diverge (in / out) along the length of the second portion 230 from the middle portion 234 of the first portion 218 to the second point 232 relative to the length of the body 210. Figure 7 and Figure 8 The page in Figure 9 (The middle one is vertical). This arrangement creates a diffuser opening of 240. In Figure 8 In another embodiment shown, the second portion 230 may diverge along the length of the second portion 230 from the middle portion 234 of the first portion 218 to the second point 220 relative to the outer surface 222 and inner surface 242 of the airfoil wall 214 to generate a diffuser opening 240. Figure 8 and Figure 9 The embodiments can exist individually or collectively. The divergent second portion 230 (and specifically, the diffuser opening 240) can be formed using any angled drilling technique currently known or developed thereafter, such as computer numerical control (CNC) milling, laser or electrical discharge machining (or electrical spark machining, i.e., EDM)).

[0056] In some embodiments, such as Figure 7 As shown, the first portion 218 may have a smaller cross-sectional area (D1) than the cross-sectional area (D2) of the second portion 230. In this way, the first portion 218 can be sized to quantitatively supply the coolant flow through the cooling channel 204. Figure 8 In another embodiment shown, the portion 238 of the second portion 230 near its entry point may have a smaller cross-sectional area (D3) than the cross-sectional area (D4) of the first portion 218. In this way, a metered flow of coolant can be supplied to the second portion 230.

[0057] Each cooling channel 204 also includes a cover element 250 that closes the first portion 218 at a first point 220, but leaves the second portion 230 open (i.e., at a second point 232). In this way, each cooling channel 204 has a single inlet 252 in fluid communication with the coolant chamber 216 and a single outlet 254 at the second point 232 of the second portion 230. That is, although the first portion 218 and the second portion 230 penetrate the outer surface 222 of the airfoil wall 214, only one outlet 254 is ultimately provided. Coolant passes through a portion of the first portion 218 and then exits through the second portion 230, and is directed downstream to form a film on the turbine airfoil 200.

[0058] The process of using the cover element 250 to close the first portion 218 at the first point 220 can take various forms. Figure 5-8In one embodiment shown, the cover element 250 comprises only a TBC 260 (e.g., an adhesive coating 262 with a TBC layer 264 or a separate TBC layer 264). Here, using the cover element 250 to enclose the first portion 218 involves forming the TBC 260 throughout the outer surface 222. The TBC 260 has sufficient thickness to enclose the first portion 218 at the first point 220. As shown, although the first portion 218 is enclosed by the TBC 260 at the first point 220, the second portion 230 of each cooling channel 204 extends through the TBC at its second point 232. The TBC 260 may naturally not enclose the second portion 230, for example, because the second point 232 is larger than the first point 220, or by employing a coating collector within the diffuser opening 240 to collect unwanted coating. Alternatively, the second portion 230 may be reopened after the TBC 260 is formed (e.g., by drilling or etching), or the second portion 230 may be machined after the outer surface 222 is coated. The TBC 260 may be shaped to fit the shape of the second portion 230 at the second point 220; for example, the TBC 260 may extend the surface of the second portion 230 to maintain the diffuser opening 240. Alternatively, the TBC 260 may be shaped differently from the second portion 230 to create a different path for the coolant. The TBC 260 may include an adhesive coating layer 262 (and, if necessary, a TBC layer 264), with the TBC layer 264 alone. The adhesive coating layer 262 may include any adhesive coating material currently known or developed thereafter (such as, but not limited to, nickel or platinum aluminum compounds, nickel-chromium-aluminum yttrium (NiCrAlY), or nickel-cobalt-chromium-aluminum yttrium (NiCoCrAlY)). TBC layer 264 may include any TBC material currently known or developed thereafter, such as, but not limited to, yttrium-stabilized zirconia (YSZ), mullite (or polyaluminate andalusite), and alumina. TBC 260 may also include additional layers (such as thermally grown oxides).

[0059] exist Figure 10In another embodiment shown in the enlarged partial cross-section, the cover element 250 may further include a metal plug 270 in the first portion 218 at or near the first point 220, for example, to fill a cavity in the outer surface 222 or to retain a small recess that can be filled by the TBC 260. Here, the use of the cover element 250 to enclose the first portion 218 includes forming the metal plug 270 in or near the first point 220 within the first portion 218. The metal plug 270 may comprise a material (e.g., metal or metal alloy or CMC) that is integrally joined (e.g., brazed or welded) to the body 210. As previously described, the TBC 260 may be formed to cover the outer surface 222 of the airfoil wall 214, and a second portion 230 of each cooling channel 204 extends through the TBC 260 at its second point 232. The TBC 260 also covers the first portion 218 and the metal plug 270.

[0060] exist Figure 11 In another embodiment shown in the enlarged partial cross-section, the cover element 250 includes a metal layer 280 that traverses the first portion 218 at a first point 220. The metal layer 280 can be bonded to the outer surface 222 of the airfoil wall 214 by brazing, welding, or other bonding methods. Here, using the cover element 250 to enclose the first portion 218 includes the metal layer 280 formed traversing the first portion 218 at the first point 220. The body 210 in this embodiment includes a metal or metal alloy. The metal layer 280 can be any suitable metal or metal alloy suitable for bonding to the body 210 at the airfoil wall 214, and will depend on the metal of the body 210. As previously described, the TBC 260 covers the outer surface 222 of the airfoil wall 214, and a second portion 230 of each cooling channel 204 extends through the TBC 260 at its second point 232. Figure 11 In this process, the metal layer 280 is not only located on the surface 222, but can also be recessed into the surface 222.

[0061] Regardless of the form of the cover element 250 used, the process may also include the preparation of the outer surface 222 to, for example, help the TBC 260 or metal layer 280 to enclose the first portion 218. The preparation process may include, but is not limited to, deformation processes (such as shot peening of the outer surface 222).

[0062] Figure 12Figure 13 illustrates another embodiment in which the inner surface 242 of the body 210 may include an enlarged inlet portion 290, to which the first portion 218 is formed. The enlarged inlet portion 290 may be formed, for example, by casting or additive manufacturing, after or together with the formation of the body 210 in any currently known manner. The enlarged inlet portion 290 can have any shape (e.g., as shown in Figure 13). Figure 12 The circle or like Figure 13 (polygons in the text).

[0063] The cooling channel 204 can be formed directly in the main body 210, or as... Figure 5 As shown, it can be provided as part of a coupon 292, which is attached to an opening in the airfoil wall 214. That is, a portion of the leading edge 212 (which includes a plurality of cooling channels 204 and a cover element 250) includes a coupon 292 attached to the remainder of the (metal) body 210. The coupon 292 can be formed using any process currently known or developed in the future (e.g., casting and drilling or additive manufacturing).

[0064] In addition to what is described elsewhere in this document, the multiple portions of the cooling channel can also have any desired cross-sectional shape (e.g., annular, elliptical, polygonal, etc.).

[0065] Embodiments of this disclosure provide turbine airfoils, blades, and / or nozzles having airfoil walls with improved cooling efficiency, allowing the turbine machinery employed therein to be more efficient or to generally reduce coolant usage through increased ignition temperatures. The two-part cooling channels described herein offer higher cooling efficiency than conventional diffuser holes, conical holes, and circular holes. The improved cooling efficiency provided can extend component life through reduced metal temperatures and a decreased likelihood of TBC delamination. Consequently, embodiments of this disclosure also reduce the likelihood of unplanned downtime and extend the duration until component maintenance is required. The surface area of ​​cooling channel 204 is larger than that of conventional holes in all embodiments, which similarly enhances internal cooling, especially when shock cooling is provided inside the coolant chamber 216. The teachings of this disclosure are particularly advantageous not only for the leading edge walls of airfoils but can also be applied to other parts of airfoils.

[0066] The terminology used herein is for the purpose of describing particular embodiments only and is not intended to limit this disclosure. As used herein, unless the context clearly indicates otherwise, the singular forms “a,” “an,” and “this” are intended to equally include the plural forms. It will be further understood that, when used in this specification, the terms “comprises and / or comprising” specify the presence of the elucidated features, integrals, steps, operations, elements, and / or components, but do not exclude the presence or addition of one or more other features, integrals, steps, operations, elements, components, and / or groups thereof. “Optional” or “optionally” means that the events or situations subsequently described may or may not occur, and the description includes instances where the event occurs and instances where the event does not occur.

[0067] As used herein throughout the specification and all claims, approximate language can be applied to modifications of any quantitative representation that can be permitted to be changed without resulting in a change in the relevant essential function. Therefore, values ​​modified with terms such as “about,” “approximately,” and “substantially” are not limited to precise specified values. In at least some instances, approximate language may correspond to the precision of the instrument used to measure the value. Here, and throughout the specification and all claims, scope limitations may be combined and / or interchanged, and unless otherwise indicated by context or language, such scopes are identified and include all subscopes contained therein. The term “approximately” applied to a specific value within a scope applies to two values, and, unless otherwise dependent on the precision of the instrument measuring the value, may indicate + / - 5% of the stated value.

[0068] All means or steps in the following claims, along with the corresponding structures, materials, actions, and equivalents of the functional elements, are intended to include any structure, material, or action for performing a function in combination with other claimed elements, as specifically claimed. This disclosure has been presented for illustrative and descriptive purposes, but is not intended to be exclusive or limited to the disclosure in the form shown. Many modifications and variations will be apparent to those skilled in the art without departing from the scope and spirit of this disclosure. Embodiments have been chosen and described to most clearly explain the principles of this disclosure and its practical application, and to allow others skilled in the art to understand this disclosure with regard to various embodiments using various modifications as suited to the particular application contemplated.

Claims

1. A turbine airfoil component, comprising: A main body with airfoil-shaped walls; The coolant chamber within the main body; and The airfoil wall contains multiple cooling channels, each including: The first part extends from a first point on the outer surface of the airfoil wall to the coolant chamber. The second part extends from a second point on the outer surface of the airfoil wall, away from the first point, to intersect with the middle portion of the first part, and A cover element that completely closes the first portion at the first point, but leaves the second portion open. Each cooling channel has a single inlet in fluid communication with the coolant chamber and a single outlet at a second point of the second portion, and the second portion meets the outer surface of the airfoil wall at the second point at an angle between 135° and 175°.

2. The turbine airfoil according to claim 1, wherein, The second point of the second part includes a diffuser opening in the axially downstream direction from the leading edge.

3. The turbine airfoil according to claim 2, wherein, The second portion radiates along its length from the middle portion of the first portion to the second point relative to the outer and inner surfaces of the airfoil wall.

4. The turbine airfoil according to claim 2, wherein, The second portion extends radially along the length of the second portion from the middle portion of the first portion to the second point at an angle between 0° and 20° relative to the length of the body.

5. The turbine airfoil according to claim 1, wherein, The second part intersects with the middle part of the first part.

6. The turbine airfoil according to claim 1, wherein, The first part meets the outer surface of the airfoil wall at the first point at an angle between 45° and 90°.

7. The turbine airfoil according to claim 1, wherein, The cover element includes a thermal barrier coating, and the second portion of each cooling channel extends through the thermal barrier coating at its second point.

8. The turbine airfoil according to claim 1, wherein, The cover element includes a metal plug at or near the first point in the first portion, and, It also includes a thermal barrier coating covering the outer surface of the airfoil wall, with the second portion of each cooling channel extending through the thermal barrier coating at its second point.

9. The turbine airfoil according to claim 1, wherein, The cover element includes a metal layer that traverses the first portion at the first point, and, It also includes a thermal barrier coating covering the outer surface of the airfoil wall, with the second portion of each cooling channel extending through the thermal barrier coating at its second point.

10. The turbine airfoil according to claim 1, wherein, A portion of the airfoil includes the plurality of cooling channels, and the cover element includes a tab that is attached to the remainder of the body.

11. The turbine airfoil according to claim 1, wherein, The cross-sectional area of ​​the first part is smaller than that of the second part, where the second part intersects with the middle part of the first part.

12. The turbine airfoil according to claim 1, wherein, The body includes a base and a tip arranged relative to the base, and pressure sidewalls and suction sidewalls extend between a leading edge and a trailing edge, wherein the airfoil wall is arranged in the leading edge.

13. A method of forming a cooling channel in the airfoil wall of an airfoil, the airfoil including a coolant chamber therein, the method comprising: A first portion forming the cooling channel extends from a first point on the outer surface of the airfoil wall to the coolant chamber; A second portion forms the cooling channel, the second portion extending from a second point on the outer surface of the airfoil wall away from the first point to intersect with the middle portion of the first portion; as well as At the first point, a cover element is used to completely close the first part, but the second part remains open. Each cooling channel has a single inlet in fluid communication with the coolant chamber and a single outlet at the second point of the second portion, and the second portion meets the outer surface of the airfoil wall at the second point at an angle between 135° and 175°.

14. The method according to claim 13, wherein, Each formation involves drilling.

15. The method according to claim 13, wherein, Forming the first portion involves drilling outward from the coolant chamber.

16. The method according to claim 13, wherein, Forming the second portion includes forming a second point to include a diffuser opening in the axially downstream direction from the leading edge.

17. The method according to claim 13, wherein, Completely enclosing the first portion using the cover element includes forming a thermal barrier coating all over the outer surface, with the second portion of each cooling channel extending through the thermal barrier coating at its second point.

18. The method according to claim 13, wherein, Using the cap element to completely close the first portion includes a metal plug forming at or near the first point in the first portion, and It also includes forming a thermal barrier coating all over the outer surface, with the second portion of each cooling channel extending through the thermal barrier coating at its second point.

19. The method according to claim 13, wherein, Using the cover element to completely enclose the first portion includes a metal layer connecting the first portion at the first point, and... It also includes forming a thermal barrier coating all over the outer surface, with the second portion of each cooling channel extending through the thermal barrier coating at its second point.

20. The method according to claim 13, wherein, Forming the first portion and the second portion includes forming a cross-sectional area of ​​the first portion that is smaller than the cross-sectional area of ​​the second portion, wherein the second portion intersects with the middle portion of the first portion.

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