Electric heating – Metal heating – By arc
Reexamination Certificate
2000-06-16
2002-08-27
Dunn, Tom (Department: 1725)
Electric heating
Metal heating
By arc
C219S121660
Reexamination Certificate
active
06441341
ABSTRACT:
BACKGROUND OF THE INVENTION
The present invention is generally directed to a method of forming cooling holes in ceramic matrix composite components, and specifically to a method of forming cooling holes in ceramic matrix composite components for use at elevated temperatures in which the matrix is a silicon carbide or a silicon nitride.
Improvements in manufacturing technology and materials are the keys to increased performance and reduced costs for many articles. As an example, continuing and often interrelated improvements in processes and materials have resulted in major increases in the performance of aircraft gas turbine engines. These improvements frequently have been in the areas of weight reduction and/or improving the temperature capabilities of the engine and its respective components, which improve engine efficiency.
An aircraft gas turbine engine or jet engine draws in and compresses air with an axial flow compressor, mixes the compressed air with fuel, burns the mixture, and expels the combustion product through an axial flow turbine that powers the compressor. The compressor includes a disk with blades projecting from its periphery. The disk turns rapidly on a shaft, and the curved blades draw in and compress air in somewhat the same manner as an electric fan. In addition to supporting combustion, the compressed air is then used to cool the engine components in the combustor and portions of the engine aft of the combustor. Additional air from the compressor is used in conjunction with auxiliary systems of the engine and plane.
Although lighter in weight than the superalloy materials typically utilized in the hot sections of gas turbine engines, ceramic matrix composites that include materials such as silicon carbide (SiC) and silicon nitride (SiN) have not been used in hot oxidizing atmospheres such as the combustor or turbine portion of gas turbine engines for various components because of problems with oxidation of SiC and SiN.
These materials have a tendency to oxidize, the SiC and SiN being converted into silica (SiO
2
), CO and CO
2
or NO
2
respectively at elevated temperatures. Furthermore, unlike metals, there has been no effective method developed for providing cooling holes to utilize cooling air to cool the component such as is done with metal components. One of the problems is that conventionally applied cooling holes provide additional surface area for the oxidation of the SiC and SiN, when used as a matrix material or when used as the fiber reinforcement. Because the holes are so small, being between about 0.010-0.030 inches, there is no effective way of applying a protective coating over the newly exposed surface area, as the methods of applying the known protective coatings could close the cooling holes, rendering them ineffective and defeating their purpose.
While various methods are available for drilling holes in turbine components used in hot portions of turbine components, these methods are primarily directed at drilling holes in metal components. Two of these methods, U.S. Pat. Nos. 4,762,464 and 4,808,785 to Vertz et al. are directed to a two step method of forming cooling holes in airfoils utilizing a combination of laser drilling and EDM. EDM is a well-know process utilizing a spark discharge in which the work tool and the workpiece are charged electrodes and the spark is a transient electric discharge through the space between the electrodes.
U.S. Pat. No. 4,873,414 to Ma et al., solves the problem of detecting when a laser breaks through a surface such as a hollow metal airfoil component by filling the hollow portion of the airfoil with a light-emitting material. U.S. Pat. No. 5,140,127 to Stroud et al. solves the problem by injecting copolymers into the cavity so that the back wall of the hollow metal airfoil is unaffected by the laser beam, which vaporizes the copolymer. U.S. Pat. No. 5,222,617 solves the same problem by laser drilling the cooling holes in the investment cast metal airfoil before removing the ceramic core utilized in the investment casting process.
U.S. Pat. No. 5,465,780 to Muntner et al. discusses a method of manufacturing complex hollow blades by using laser machining for forming an intricate ceramic core. After casting the blade by pouring alloy around the core, the core is removed by conventional leaching methods.
U.S. Pat. No. 5,683,600 to Kelley et al. sets forth a method of drilling compound cooling holes with a non-circular surface opening in a gas turbine engine with a laser beam in metal alloys such as steels, titanium alloys, inconels and other nickel based superalloys. The method overcomes the conductive and reflective nature of these superalloys which causes a waveguiding effect to occur on the laser beam. The method controls the focal spot below or to undershoot the surface a preselected distance D to overcome the problems with the prior art processes and allows for the formation of the complex hole without the need to resort to the additional steps of EDM. U.S. Pat. No. 5,837,964 to Emer et al. sets forth a process for laser drilling large and deep holes in superalloy components by utilizing a combination of laser drilling a small central hole followed by trepanning laser drilling the hole to final size.
However, all of the prior art processes are directed to the problems with drilling laser holes for cooling, in metallic, typically superalloy, components. None of the prior art processes describes the problems associated with drilling small cooling, holes in oxidizable, non-metallic components for use in the hot section of gas turbine engines, such as SiC or SiN-containing ceramic matrix composites (CMCs). Thus, none of the prior art processes sets forth solutions to these problems which include drilling holes to an adequate size to permit cooling air to flow through the non-metallic components without oxidizing the oxidizible component and without cracking the relatively brittle ceramic material, while also providing a protective coating over the newly created surface area of the hole, so that the air flowing through the holes does not oxidize the oxidizable component of the CMC.
SUMMARY OF THE INVENTION
A method for producing apertures in turbine airfoil components and combustor liners utilizes laser drilling of ceramic matrix composites that have at least one oxidizable component. Air-cooled CMCs have not been used commercially in the hot section of gas turbine engines, that is, the portion of the engine that includes the combustor, as these materials exhibit deficiencies, despite the weight advantage that these materials have over the typical metallic superalloy components. As these deficiencies are overcome, the prospects of these materials being used increase. One of these deficiencies has been an inability to drill small holes in CMCs having an oxidizable component while providing suitable protection to the newly formed surface of the cooling holes so that the oxidizable component will not decompose, contributing to part failure, as cooling air is circulated through them.
Examples of turbine components that can be made from CMCs utilzing the parameters of the present invention include turbine blades, turbine vanes, turbine buckets, nozzles, and the like. Some combustor components such as combustorliners can also be manufactured. The parameters developed for laser drilling of the holes produces holes of predetermined size, geometry and hole pattern to promote effective cooling. As the laser is applied to the ceramic matrix composite material, a portion of the material is vaporized or ablated to form the aperture or hole. However, some of the energy of the laser beam also melts material adjacent to the beam. This melted material is oxidized and briefly flows along the newly formed aperture surface. This material quickly forms a recast layer along the aperture surface as it cools which forms an oxidation barrier for the oxidizable ceramic matrix material under it. This oxidation barrier prevents or at least significantly reduces the oxidation of the underlying CMC, thereby reducing or eliminating
Edmondson Wayne Garcia
Scheidt Wilbur Douglas
Steibel James Dale
Dunn Tom
General Electric Company
Johnson Jonathan
Maria Carmen Santa
McNees Wallace & Nurick L.
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