Patent Application
20050102835
This application relates generally to gas turbine engines and, more particularly, to methods for repairing gas turbine engine rotor blades.
At least some known gas turbine engines include a compressor for compressing air which is mixed with a fuel and channeled to a combustor wherein the mixture is ignited within a combustion chamber for generating hot combustion gases. The hot combustion gases are channeled downstream to a turbine, which extracts energy from the combustion gases for powering the compressor, as well as producing useful work to propel an aircraft in flight or to power a load, such as an electrical generator.
Known compressors include a rotor assembly that includes at least one row of circumferentially spaced rotor blades. Each rotor blade includes an airfoil that includes a pressure side, and a suction side connected together at leading and trailing edges. Each airfoil extends radially outward from a rotor blade platform. Each rotor blade also includes a dovetail that extends radially inward from a shank coupled to the platform. The dovetail is used to mount the rotor blade within the rotor assembly to a rotor disk or spool. In at least some known compressors, the rotor blade is formed integrally with the rotor disk or spool.
During operation, leading and trailing edges of the blade and/or a tip of the compressor blade may deteriorate or become damaged due to any of a number of distress modes, including, but not limited to, foreign object damage (FOD), tip rubbing, oxidation, thermal fatigue cracking, or erosion caused by abrasives and corrosives in the flowing gas stream. To facilitate mitigating such operational effects, the blades are periodically inspected for damage, and a determination of an amount of damage and/or deterioration is made. If the blades have lost a substantial quantity of material they are replaced. If the blades have only lost a small quantity material, they may be returned to service without repair. Alternatively, if the blades have lost an intermediate quantity of material, the blades may be repaired.
For example, at least one known method of repairing a turbine compressor blade includes mechanically removing, such as by grinding, a worn and/or damaged tip area and then adding a material deposit to the tip to form the tip to a desired dimension. The material deposit may be formed by several processes including welding and/or thermal spraying. Furthermore, special tooling is also used to achieve the precise dimensional relations between the original portion of the compressor blade and the added portion of the compressor blade. Thus, replacing a portion of a compressor blade may be a time-consuming and expensive process. Additionally, more complex airfoil shapes, for example three-dimensional aerodynamic configurations may increase the difficulty of welding and blending the repaired blade, thus resulting in increased repair costs.
In one aspect, a method for replacing a portion of a gas turbine engine rotor blade, the rotor blade having a contour defined by a blade first sidewall and a blade second sidewall is provided. The method includes cutting through the rotor blade such that a cut line extends from a leading edge of the blade to a trailing edge of the blade, and between the first sidewall and the second sidewall, removing the portion of the rotor blade that is radially outward of the cut line, and coupling a replacement blade portion to remaining blade portion such that the newly formed rotor blade is formed with a pre-determined aerodynamic contour.
In another aspect, a method for replacing a portion of a gas turbine engine rotor blade including a leading edge, a trailing edge, a first sidewall, and a second sidewall, and having a contour defined by the first sidewall and the second sidewall is provided. The method includes uncoupling the rotor blade from the gas turbine engine, cutting through the rotor blade such that a cut line extends from the leading edge to the trailing edge, and between the first sidewall and the second sidewall, removing the portion of the rotor blade radially outward of the cut line, coupling a replacement blade portion to the remaining blade portion, and contouring the replacement blade portion such that the newly formed rotor blade is formed with a pre-determined aerodynamic contour.
In a further aspect, a method for replacing a portion of a gas turbine engine rotor blade including a leading edge, a trailing edge, a first sidewall, and a second sidewall, and having a contour defined by the first sidewall and the second sidewall is provided. The method includes uncoupling a compressor rotor blade from the gas turbine engine, cutting through a portion of the rotor blade such that a cut line extends from the leading edge to the trailing edge, and between the first sidewall and the second sidewall, removing a portion of the rotor blade radially outward of the cut line, welding a replacement blade portion to the portion of the compressor rotor blade remaining, and contouring the replacement blade portion such that the newly formed compressor rotor blade has a contour that substantially mirrors that of the original compressor rotor blade contour.
In operation, air flows through fan assembly 12 and compressed air is supplied to high pressure compressor 14 through booster 22. The highly compressed air is delivered to combustor 16. Airflow (not shown in
Each airfoil 60 includes a first sidewall 70 and a second sidewall 72. First sidewall 70 is convex and defines a suction side of airfoil 70, and second sidewall 72 is concave and defines a pressure side of airfoil 60. Sidewalls 70 and 72 are joined at a leading edge 74 and at an axially-spaced trailing edge 76 of airfoil 60. More specifically, airfoil trailing edge 76 is spaced chord-wise and downstream from airfoil leading edge 74.
First and second sidewalls 70 and 72, respectively, extend longitudinally or radially outward in span from a blade root 78 positioned adjacent platform 62, to an airfoil tip 80. Airfoil tip 80 defines a radially outer boundary of an internal cooling chamber 82. Cooling chamber 82 is bounded within airfoil 60 between sidewalls 70 and 72, and extends through platform 62 and through shank 64 and into dovetail 66. More specifically, airfoil 60 includes an inner surface 83 and an outer surface 84, and cooling chamber 82 is defined by airfoil inner surface 83.
Platform 62 extends between airfoil 60 and shank 64 such that each airfoil 60 extends radially outward from each respective platform 62. Shank 64 extends radially inwardly from platform 62 to dovetail 66. Dovetail 66 extends radially inwardly from shank 64 and facilitates securing rotor blade 50 to rotor disk 26.
Deteriorated and/or damaged regions 86 of rotor blade 50 may be removed and replaced using the methods described herein. More specifically, deteriorated and/or damaged regions 86 of airfoil 60 including leading edge 74, trailing edge 76, and airfoil tip 80, may be removed and replaced using the methods described herein. If an engine, such as engine 10, indicates that rotor blade 50 includes at least one damaged and/or deteriorated portion 86 of rotor blade 50 is removed from engine 10 and repaired using the methods described herein.
More specifically, as shown in
In one embodiment, a joint 152 between replacement tip 120 and preserved portion 98 may be configured and placed where it can be a simple geometry, and then welded using a high yield automated process. Additionally, undamaged portion 120 may be fabricated from a material similar to damaged portion 90 thereby more closely matching the original material, i.e. forged vs. cast. In the exemplary embodiment, the methods described herein can be adapted to weld common blade alloys such as, but not limited to, a nickel based alloy, a titanium based alloy, and an iron based alloy, i.e. A286. Additionally, the methods described herein provides superior weld properties and facilitates improving control of the airfoil shape and orientation, while reducing distortion compared to other known compressor blade repair methods. Further, a single weld joint facilitates reducing weld defects since other known methods require multiple pass welding material build up. Accordingly, there is less weld area to fluorescent penetrant inspect or X-ray using the resistance projection weld methods described herein.
Although the repair methods described herein are described in the context of a compressor blade, it should be realized that the methods described herein are equally applicable to turbine rotor blades, power turbine rotor blades, low pressure compressor rotor blades, and fan rotor blades. The repair methods can also be used to repair fan, compressor, or turbine stators if their configuration allows removal of a damaged portion of the stator airfoil.
The above-described airfoil repair methods enable an airfoil having damage and/or deterioration extending along its leading and/or trailing edges, and/or along its airfoil tip, to be repaired in a cost-effective and reliable manner. More specifically, the above-described airfoil repair methods facilitate restoring a damaged and/or deteriorated blade to its original dimensions. Accordingly, using the methods described, the entire top end of the blade is removed. A portion of blade having the same contour as the original blade contour is welded back to the salvaged part of the blade. The repair methods described herein offer a plurality of advantages over known methods. Specifically, turbine engine 10 is returned to service using a repair process that facilitates improved savings in comparison to removing and replacing entire turbine blades, or alternatively adding weld filler metal to the blade tip to build up the tip to a desired dimension.
Exemplary embodiments of blade repair methods are described above in detail. The repair methods are not limited to the specific embodiments described herein, but rather, components and aspects of each repair method may be performed and utilized independently and separately from other repair methods described herein. Moreover, the above-described repair methods can also be used in combination with other repair methods and with other rotor blade or stator components. Specifically, the above-described repair methods can also be used to repair bladed disks, i.e. blisks, integrated disks, and blades in a single component.
While the invention has been described in terms of various embodiments, those skilled in the art will recognize that the invention can be with modification within the spirit and scope of the claims.
