Air-cooled gas turbine components and methods for their manufacture and repair

Abstract

Method for manufacturing a component suitable for use in a gas turbine engine (52), the method comprising the steps: Generating the component (52) from a substrate having a first surface and a second surface (33); Creating at least one cooling hole (160) through the substrate from the first surface to the second surface (33), wherein the cooling hole (160) has a first opening cross-section and an axis (220) which is oriented at a non-perpendicular angle (α) to the second surface (33); Applying a first coating (212) to at least one of the first surface and the second surface (33) adjacent to the at least one cooling hole (160), wherein the cooling hole (160) remains at least partially uncovered by the first coating (212); Applying a second coating (214) to the first coating (212) adjacent to the at least one cooling hole (160), wherein the cooling hole (160) remains at least partially uncovered by the second coating (214), wherein the first and second coatings (212, 214) initially form a small opening with a second opening cross-section (162) and with an axis (230) that is oriented at a larger angle (β) to the second surface (33) than the axis (220) of the cooling hole (160); and Removing the second coating (214) from the cooling hole (160), leaving most or all of the first coating (212) intact, to define the second opening cross-section (162), which is smaller than the first opening cross-section, and to reduce the angle (β) of the axis (230) of the small opening having the second opening cross-section (162).

Description

BACKGROUND OF THE INVENTION

[0001] The technology described herein relates primarily to gas turbine engines, and more particularly to air-cooled components for use in gas turbines and methods for manufacturing and repairing such components.

[0002] A gas turbine engine includes a compressor for compressing air, which is mixed with fuel and fed to a combustion chamber, where the mixture is ignited within a combustion chamber to produce hot combustion gases. At least some known combustors include a dome assembly, a nozzle, and inserts for directing the combustion gases to a turbine, which extracts energy from the combustion gases to drive the compressor and generates useful work for propelling an aircraft in flight or driving a load, such as an electrical generator. The inserts are connected to the dome assembly by means of the nozzle and extend downstream from the nozzle to define the combustion chamber.

[0003] The operating environment within a gas turbine engine is both thermally and chemically aggressive. Significant advances in high-temperature alloys have been achieved through the formulation of iron-, nickel-, and cobalt-based superalloys. Yet, components made from such alloys often cannot withstand long-term service loads when used in certain sections of a gas turbine engine, such as the turbine, combustor, or afterburner. A common solution is to protect the surfaces of such components with an environmental coating system, such as an aluminide coating or a thermal-barrier-coating (TBC) system. The latter typically comprises an environmentally resistant bond coat and a ceramic thermal-barrier coating deposited on the bond coat.Bond layers are typically made from an oxidation-resistant alloy, such as McrAlY, where M is iron, cobalt and / or nickel, or from a diffusion aluminide or platinum aluminide, which form an oxidation-resistant intermetallic compound.

[0004] Although thermal barrier coating systems provide significant thermal protection for the underlying component substrate, internal cooling of components, such as combustor liners, is generally required and can be used in combination with or instead of a thermal barrier coating. Gas turbine engine combustor liners often require a complex cooling scheme in which cooling air flows around the combustion chamber and is then discharged into the combustion chamber through carefully configured cooling holes in the combustor liner. The performance of a combustor is directly related to the ability to achieve uniform cooling of its surfaces with a limited amount of cooling air.Consequently, processes by which cooling holes and their openings are created and configured are often critical, as the size and shape of each opening determine the amount of airflow exiting the opening and the distribution of airflow over the surface, affecting the overall flow distribution within the combustion chamber. Other factors, such as the local surface temperature of the insert, are also influenced by variations in opening size.

[0005] For combustion chamber liners with a thermal barrier coating, cooling holes are typically created using conventional drilling techniques such as electric discharge machining (EDM) and laser machining. However, EDM cannot be used to create cooling holes in a combustion chamber liner with a ceramic TBC because the ceramic is electrically non-conductive, and laser machining is prone to spalling the brittle ceramic TBC by cracking at the interface between the substrate and the ceramic. Consequently, it has been necessary to create cooling holes using EDM and / or laser machining prior to applying the TBC system, limiting the thickness of the TBC that can be applied or requiring final machining to remove ceramic from the cooling holes to restore the desired size and shape of the openings.Conventional processes involve protecting cooling holes from TBC deposition or completely removing deposited TBC from the holes to achieve the desired hole geometry. This results in exposure of the underlying metal surface to the aggressive environmental conditions at the hole locations.

[0006] Current repair methods for air-cooled components, such as combustion chamber liners, involve welding thermal fatigue cracks. The location of openings in the panels, such as cooling or dilution holes, and the use of thermal barrier coatings introduce additional complexity for the application of welds and inserts. In many cases, protective coatings must be removed from an entire panel and / or insert to access the underlying metal itself, and then the protective coatings reapplied. However, conventional reapplication processes involve protecting cooling holes from TBC deposition or completely removing applied TBC from the holes to achieve the desired hole geometry. This results in the underlying metal surface being exposed to the aggressive environmental conditions at the hole locations.In some cases, repairing such panels is not an option and instead the entire combustion chamber insert must be replaced.

[0007] Because conventional designs may utilize the underlying metal substrate to define the final hole geometry without a TBC system applied to the hole surfaces, damage to the holes or repairs made to the holes provided in the metal substrate can compromise the performance of the repaired part. Additionally, a process for manufacturing air-cooled components, such as combustor liners, is desired in a manner that is economically and physically feasible, provides improved protection for the substrate near the cooling holes, and results in satisfactory cooling hole geometry during both manufacture and repair.

[0008] US 6,210,488 B1 describes a method for removing a thermal barrier coating from cooling holes in a surface of a component suitable for use in a gas turbine power plant. The thermal barrier coating is removed both from the surface and in the area of ​​a hole, and ultrasonic energy is then directed into the hole.

[0009] A method for adjusting an air flow through a plurality of cooling holes is described in US Pat. No. 6,408,610 B1. In one embodiment, a component having the cooling holes comprises two layers, with the outer layer covering the inner layer up to the cooling hole.

[0010] A protective coating which extends into a cooling hole in a component suitable for a gas turbine power plant is also known from US 5 941 686 A.

[0011] WO 2006 / 106061 A1 relates to a method for repairing or renewing cooling holes in a coated component of a gas turbine. This involves applying new coatings to a damaged area around a cooling hole, which then initially partially or completely seal the cooling hole. Subsequently, the entire cooling hole cross-section is exposed again in the coatings.

[0012] From US 2004 / 0 134 897 A1 it is known to coat a component with a hole and then to expose the hole in the coating that was initially closed by the coating.

[0013] US 6,329,015 B1 relates to a method for forming shaped holes in a substrate coated with a ceramic layer. After the hole is created, an abrasive fluid stream is introduced into the hole to open its cross-section and achieve the desired shape. SUMMARY OF THE INVENTION

[0014] The invention relates to a method for producing a component suitable for use in a gas turbine engine according to patent claim 1 and to a method for repairing a component suitable for use in a gas turbine engine according to patent claim 5. BRIEF DESCRIPTION OF THE DRAWINGS

[0015] The accompanying drawings illustrate various embodiments of the technology described herein, wherein: Fig. 1 is a schematic diagram of an exemplary gas turbine engine; Fig. 2 is a schematic cross-sectional view of an exemplary combustion chamber arrangement used in the Fig. 1; Fig. 3 is an enlarged perspective view of a portion of an exemplary combustor liner used in the Fig. 2 shown combustion chamber arrangement; Fig. 4 is an enlarged partial cross-sectional view of the Fig. 3 shown combustion chamber insert before coating application; and Fig. 5 is an enlarged partial cross-sectional view of the Fig. 4 shown combustion chamber insert after a coating application; Fig. 6 is an enlarged partial cross-sectional view of the Fig. 5 shown combustion chamber insert after removal of part of the coating material; Fig. 7 is a flowchart illustrating the steps associated with an exemplary manufacturing process; and Fig. Figure 8 is a flowchart illustrating the steps associated with an exemplary repair procedure. DETAILED DESCRIPTION OF THE INVENTION

[0016] The present invention is generally applicable to air-cooled components, and particularly to those protected from a thermally and chemically aggressive environment by a thermal barrier coating system. Notable examples of such components include the high- and low-pressure turbine nozzles and blades, shrouds, combustor liners, and afterburner hardware of gas turbine engines. The advantages of this invention are particularly applicable to gas turbine engine components that utilize internal cooling and a thermal barrier coating to maintain the component's operating temperature at an acceptable level while operating in a thermally aggressive environment.

[0017] Fig. 1 is a schematic representation of an exemplary gas turbine engine 10. The engine 10 includes a low-pressure compressor 12, a high-pressure compressor 14, and a combustor assembly 16. The engine 10 also includes a high-pressure turbine 18 and a low-pressure turbine 20 arranged in series, axial flow relationship. The compressor 12 and the turbine 20 are coupled by a first shaft 21, and the compressor 14 and the turbine 18 are coupled by a second shaft 22. In the exemplary embodiment, the gas turbine engine 10 is a CFM-56 engine, commercially available from CFM International, Inc., Cincinnati, Ohio. In another embodiment, the gas turbine engine 10 is a CF-34 engine, commercially available from GE's Aviation Business, Cincinnati, Ohio.

[0018] During operation, air flows through the low-pressure compressor 12, and compressed air is supplied from the low-pressure compressor 12 to the high-pressure compressor 14. The highly compressed air is supplied to the combustion chamber 16. The air flow from the combustion chamber 16 drives the turbines 18 and 20 and exits the gas turbine engine 10 via a nozzle (not labeled).

[0019] Fig. Figure 2 is a schematic cross-sectional view of an exemplary combustion chamber 16 used in the (in Fig. 1). The combustor 16 includes an outer liner 52 and an inner liner 54 disposed between an outer combustor shell 56 and an inner combustor shell 58. The outer and inner liners 52 and 54 are radially spaced from one another to define a combustion chamber 60 therebetween. The outer liner 52 and the outer shell 56 form an outer passage 62 therebetween, and the inner liner 54 and the inner shell 58 form an inner passage 64 therebetween. A baffle assembly 66 is connected to the upstream ends of the outer and inner liners 52 and 54, respectively. An annular opening 68 formed in the baffle assembly 66 allows compressed air to enter the combustor 16 through a diffused opening in a direction generally shown by arrow A.The compressed air flows through the annular opening 68 to assist combustion and enable cooling of the inserts 52 and 54.

[0020] An annular dome plate 70 extends between and is connected to the outer and inner liners 52 and 54 near their upstream ends. A plurality of circumferentially spaced swirl assemblies 72 are connected to the dome plate 70. Each swirl assembly 72 receives compressed air from the port 68 and fuel from a corresponding fuel injector 74. The fuel and air are swirled and mixed by the swirl assembly 72, and the resulting fuel / air mixture is discharged into the combustion chamber 16. The combustion chamber 16 includes a longitudinal axis 75 extending from a forward end 76 to a rearward end 68 of the combustion chamber. In the exemplary embodiment, the combustion chamber 16 is a single annular combustion chamber. Alternatively, the combustion chamber 16 may be any other combustion chamber, including, but not limited to, a double annular combustion chamber.

[0021] In the exemplary embodiment, the outer and inner inserts 52 and 54 each include a plurality of overlapping panels 80. More specifically, in the exemplary embodiment, the outer insert 52 includes five panels 80 and the inner insert 54 includes four panels 80. In an alternative embodiment, both the outer and inner inserts 52 and 54 may each include any number of panels 80. The panels 80 define a combustion space 60 within the combustion chamber 16. More specifically, in the first embodiment, a first pair of panels 82 positioned upstream define a primary combustion zone 84, a pair of second panels 86 positioned downstream of the first panels 82, an intermediate combustion zone 88, and a pair of third panels 90 positioned downstream of the second panels 86 (direction B in Fig. 3) and a pair of fourth panels 92 positioned downstream of the third panels 90, a downstream dilution combustion zone 94.

[0022] The combustion chamber liners may include dilution holes to supply air into the combustion environment within the combustion chamber to alter the temperature distribution or combustion characteristics. Dilution air is primarily introduced into the combustion arm 60 through a plurality of circumferentially spaced dilution holes 96 extending through either one or both of the outer and inner liners 52 and 54. In the exemplary embodiment, the dilution holes 96 are each substantially circular. The dilution holes may be customized (sized, shaped, and / or arranged) as needed to meet durability and performance specifications of the particular component and product application.

[0023] Fig. 3 illustrates an exemplary combustor liner 52 that may be used with the combustor 16. The liner 52 may also include a plurality of cooling holes 160 formed in the third panel 90 to facilitate cooling of the liner 52. Although only one group of cooling holes 160 is illustrated in the third panel 90, it should be understood that the group of cooling holes 160 is spaced circumferentially around the third panel 90. It should be appreciated that each group of cooling holes 160 is positioned corresponding to particularly hot spots to facilitate channeling of the cooling fluid to the corresponding hot spot. The third panel 90 includes any number of cooling holes 160 to facilitate cooling of the liner 52.

[0024] During operation of a gas turbine 10, an inner surface 33 of the insert 52 becomes hot and requires cooling. Accordingly, in the exemplary embodiment, cooling features, such as cooling holes 160, are positioned in the insert 52 to facilitate the directing of cooling fluid to the hot spots of the insert 52. In particular, the cooling holes 160 direct cooling fluid from an outer channel 62 and / or inner channel 64 to the combustion chamber 60, thus creating a layer of cooling fluid on the inner surface 33. It should be appreciated that other embodiments may utilize any configuration of cooling holes 160 that enables the cooling holes 160 to function as described herein. Likewise, cooling holes 160 could be located in the insert 54 to cool its outer surface.

[0025] During operation, as atomized fuel is injected into the combustion chamber 60 and ignited, heat is generated in the combustion chamber 60. Although air enters the combustion chamber 60 through the cooling devices 160 and forms a thin protective boundary layer of air along the combustor liner surface 33, a change in the exposure of the combustor liner surfaces to high temperatures can induce thermal stresses in the panels 80. As a result of continued exposure to thermal stress, the panels 80 can deteriorate over time.

[0026] Fig. 4 is an enlarged partial cross-sectional view of a portion of the combustion chamber liner 52 to illustrate the relationship between the cooling holes 160 and the liner surface 33 and the axis 220 of the hole 160.

[0027] According to the Fig. 5 and Fig. 6, a layer 210 of a thermal barrier material is applied to the combustion chamber insert surface 33 of the Fig. 4. The thermal barrier material insulates the combustion chamber liner surface from high-temperature combustion gases. Layer 210 includes an inner layer 212, such as an adhesion layer, and an outer layer 214, such as a thermal barrier layer.

[0028] The exemplary methods are described with respect to an air-cooled component, such as a combustor liner 52, whose metallic substrate 33 is protected by a thermal barrier coating system consisting of an adhesion layer 212 formed on the substrate (inner surface 33) and a ceramic layer 214 adhered to the surface 33 with the adhesion layer 212. The adhesion layer 212 and ceramic layer 214 may each be a single layer of material or be constructed of two or more layers (i.e., multi-layered) of suitable materials. As is the case with high-temperature components of a gas turbine engine, the surface 33 may be an iron-, nickel-, or cobalt-based superalloy.The bond layer 212 is preferably an oxidation-resistant composition, such as a diffusion aluminide or MCrAlY, which forms an aluminum oxide (Al2O3) layer or scale (not shown) on its surface during exposure to elevated temperatures. The aluminum oxide scale protects the underlying superalloy surface 33 from oxidation and provides a surface to which the ceramic layer 214 adheres more tenaciously.

[0029] The ceramic layer 214 can be deposited using air plasma spraying (APS), low-pressure plasma spraying (LPPS), or physical vapor deposition (PVD) techniques, such as electron beam physical vapor deposition (EBPVD), the latter resulting in a strain-tolerant columnar grain structure. An exemplary material for the ceramic layer 214 is zirconium oxide partially stabilized with yttria (yttria-stabilized zirconium, or YSZ), although fully yttria-stabilized zirconium oxide could also be used, as could other oxides such as magnesium oxide (MgO), calcium oxide (CaO), cerium oxide (CeO2), or scandium oxide (Sc2O3).

[0030] The process of this invention involves the preparation of a (in the Fig. 4 to 6) which can protrude through the ceramic layer 214, the adhesive layer 212 and the surface 33 via an opening 162 to achieve a configuration for the cooling hole 160 and an opening 162 that creates a suitably metered distribution of cooling air over the outer surface of the component, such as the insert 52. As shown in Fig. 5, the cooling hole opening 162, as initially coated, forms a small opening (with axis 230) at a steep angle (angle β) to the surface. As shown in Fig. 6, after removal of the portion of the ceramic layer 214 aligned with the hole, the opening 162 is at a relatively shallow angle to the surface 33 so that the cooling air flowing through the opening 162 can form an effective film over the component surface during operation. Fig. 7 and Fig. 8 illustrate in flowchart form the exemplary methods described in more detail herein. Although both methods share some common steps, method 200 is particularly suited for the manufacture of new air-cooled components, while method 300 is particularly suited for the repair and restoration of air-cooled components during their operational lifetime.

[0031] According to the presentation in Fig. 4, a first step of this exemplary method is to create a hole through the insert 52. A second step is then the application of the adhesive layer 212 and the ceramic layer 214 on the surface 33 as shown in Fig. 5. Due to the coating buildup on the edges of the hole 160, the resulting hole opening 162 is smaller in cross-sectional diameter than the cooling hole 160 required for the insert 52, but is not completely obscured, so that the location of the hole and at least a portion of its cross-section are substantially unobstructed. For example, for a cooling hole 160 having a diameter of about 0.9 mm (0.035 inches) to about 1.0 mm (0.04 inches), the opening 162 after coating preferably has a diameter of about 0.5 mm (0.020 inches), or about half the intended diameter for the cooling hole 160, so that a "detection hole" remains visible and accessible through the coatings.Suitable techniques for creating hole 160 include EDM, although it is foreseeable that hole 160 could be created using other methods, such as casting, laser cutting, or drilling with an abrasive water jet. As a result of the drilling process, hole 160 has a substantially uniform circular cross-section and forms a non-perpendicular angle (angle α) to surface 33.

[0032] Once the hole 160 is formed and the bonding layer 212 and the ceramic layer 214 are applied, the component (the insert 52) ​​is machined using a carefully controlled process that applies a pressurized fluid stream directed at the hole 160, such as from the uncoated side of the insert 523, so as to Fig.5 and the opening 162. Various fluids could be used, such as air or water containing a medium such as glass beads or polishing sand to exert an abrasive action on the coating materials above the hole 160.

[0033] It has been found that an operation as described herein applies sufficient energy to enlarge the opening 162 to the desired dimension and angle by removing the ceramic TBC layer, but not the bond coat or the underlying starting material, such as the metal substrate. Therefore, although the operation removes the ceramic layer 214, most or all of the underlying bond coat 212 remains on the surface of the opening adjacent the cooling hole 160, so that the bond coat provides protection for the edges of the insert near the cooling hole both during manufacture and in service. Because the operation applies mechanical energy rather than thermal energy, it does not damage or chip the bond coat 212 or ceramic layer 214 surrounding the hole 160 and forming the edges of the resulting hole opening 162.

[0034] The process is capable of properly sizing and creating cooling holes and openings through a ceramic thermal barrier coating (TBC) and its underlying substrate. The abrasive fluid flow also serves to finish the hole and its opening, including the desired size and shape of the hole and opening, without removing or damaging the ceramic surrounding the cooling hole and opening.

[0035] When a returned-from-use engine, such as engine 10, reveals that the combustor liner 52 contains at least one damaged panel 80, a variety of repair methods may be employed to restore the combustor liner 52 to a serviceable condition. These repair methods may include replacing the entire liner, a complete panel, and / or a section or segment of a liner panel, as well as repairing cracks, such as by sealing them with a weld.

[0036] During a repair operation, all dirt, foreign material, and coatings are typically removed from a component, such as a combustion chamber liner, to allow for detailed inspection of the component. Any defects in the substrate, such as cracks, are then repaired using appropriate and approved methods, such as welding, brazing, or replacing discrete sections of the component. Holes, such as cooling holes, can be reamed and / or repaired as needed to restore them to the correct size, shape, and pattern.

[0037] Once the component surfaces have been suitably repaired, protective thermal barrier coatings can be applied to the component surfaces using the process described above. Because the finished opening dimensions are carefully controlled and defined by a removable and replaceable coating system as described herein, it is possible to perform and repeat the repair process while maintaining the dimensions of the finished cooling holes within specifications.

[0038] Because components, such as damaged liners, are repaired using the method described herein using readily available coating techniques, combustors can be returned to service using a repair process that provides improved savings compared to removing and replacing entire combustor liners or large patches or entire panels.

[0039] Although the devices and methods described herein have been described in the context of cooling holes in a combustor liner of a gas turbine engine, it should be understood that the devices and methods are not limited to gas turbine engines, combustor liners, or cooling holes. Likewise, the illustrated gas turbine engine and combustor liner components are not limited to the specific embodiments described herein; rather, components of both the gas turbine engine and combustor liner may be used independently and separately from other components described herein.

[0040] Although the invention has been described with respect to various specific embodiments, it will be apparent to those skilled in the art that the invention may be practiced with modifications within the spirit and scope of the claims.

Claims

[1] A method of manufacturing a component (52) suitable for use in a gas turbine engine, the method comprising the steps of: producing the component (52) from a substrate having a first surface and a second surface (33); Creating at least one cooling hole (160) through the substrate from the first surface to the second surface (33), the cooling hole (160) having a first opening cross-section and an axis (220) oriented at a non-perpendicular angle (α) to the second surface (33); Applying a first coating (212) to the at least one of the first surface and the second surface (33) adjacent to the at least one cooling hole (160), wherein the cooling hole (160) remains at least partially uncovered by the first coating (212); Applying a second coating (214) to the first coating (212) adjacent to the at least one cooling hole (160), wherein the cooling hole (160) remains at least partially uncovered by the second coating (214), wherein the first and second coatings (212, 214) initially form a small opening with a second opening cross-section (162) and with an axis (230) oriented at a greater angle (β) to the second surface (33) than the axis (220) of the cooling hole (160); and Removing the second coating (214) from the cooling hole (160), leaving most or all of the first coating (212) to define the second opening cross-section (162) which is smaller than the first opening cross-section, and to reduce the angle (β) of the axis (230) of the small opening having the second opening cross-section (162). [2] The method of claim 1, wherein at least one of the removing steps is performed by a stream of abrasive media. [3] The method of claim 2, wherein the stream of abrasive media comprises glass beads suspended in an air stream. [4] The method of claim 2, wherein the flow of abrasive media is directed through the at least one cooling hole (160) from a non-coated side of the substrate. [5] A method of repairing a component (52) suitable for use in a gas turbine engine, the component having a substrate with first and second surfaces (33) and at least one cooling hole (160) extending through the substrate from the first surface to the second surface (33), the cooling hole (160) having a first opening cross-section and an axis (220) oriented at a non-perpendicular angle (α) to the second surface (33), the method comprising the steps of: removing coatings from the component (52); repairing any defects in the substrate of the component (52); Applying a first coating (212) to at least one of the first surface and the second surface (33) adjacent to the at least one cooling hole (160), wherein the cooling hole (160) remains at least partially uncovered by the first coating (212); Applying a second coating (214) to the first coating (212) adjacent to the at least one cooling hole (160), wherein the cooling hole (160) remains at least partially uncovered by the second coating (214), wherein the first and second coatings (212, 214) initially form a small opening with a second opening cross-section (162) and with an axis (230) oriented at a greater angle (β) to the second surface (33) than the axis (220) of the cooling hole (160); and Removing the second coating (214) from the cooling hole (160), leaving most or all of the first coating (212) in place, to define a second opening cross-section that is smaller than the first opening cross-section, and to reduce the angle (β) of the axis (230) of the small opening having the second opening cross-section (162). [6] The method of claim 5, wherein at least one of the removing steps is performed by a stream of abrasive media. [7] The method of claim 6, wherein the stream of abrasive media comprises glass beads suspended in an air stream. [8] The method of claim 6, wherein the flow of abrasive media is directed through the at least one cooling hole (160) from a non-coated side of the substrate.

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