Patent Application
20090313823
This invention relates to gas turbine engine airfoil edge repair and, in particular, cutting out a damaged area and welding in beads of material to build up airfoil leading and trailing edges and tips.
Gas turbine engines include fan, compressor, combustion, and turbine sections. Disposed within the fan, compressor, and turbine sections are alternating annular stages of circumferentially disposed moving blades and stationary vanes having airfoils with leading and trailing edges and radially outer tips subject to wear and tear. The rows or stages of vanes and blades are concentrically located about a centerline axis of the gas turbine engine. The blades are typically mounted on a disk which rotates about its central axis and integrally formed disks and blades referred to as BLISKS have been used in many aircraft gas turbine engines.
Fan and compressor blades are typically forged from superalloys such as a nickel-base alloy while turbine blades typically are made from high temperature alloys or superalloys containing titanium. In addition, the casting of turbine vanes and blades is frequently performed so as to produce a directionally solidified part, with grains aligned parallel to the axis of the blade or a single crystal part, with no grain boundaries. More recently, ceramic matrix composite and metal matrix composite materials have been used to make solid and hollow gas turbine engine blades and vanes.
In service, damage and deterioration of leading and trailing edges and tip of the compressor blade occurs due to oxidation, thermal fatigue cracking and metal erosion caused by abrasives and corrosives in the flowing gas stream as well as high cycle fatigue (HCF). During periodic engine overhauls, the blades are inspected for physical damage and measurements are made to determine the degree of deterioration and damage. If the blades have lost substantial material, then they are replaced or repaired.
Several methods exist for repairing the worn or damaged turbine blades and vanes. Repair methods include, for example, conventional fusion welding, plasma spray as described in U.S. Pat. No. 4,878,953, and the use of a tape or slurry material containing a mixture of a binder and a metal alloy powder which is compatible with the substrate alloy. U.S. Pat. No. 4,878,953 provides an excellent source of background information related to methods for refurbishing cast gas turbine engine components and, particularly, for components made with nickel-base and cobalt-base superalloys for use in the hot sections of gas turbine engines and, more particularly, for components exposed to high temperature operating conditions. U.S. Pat. No. 4,726,104, entitled “Methods for Weld Repairing Hollow, Air Cooled Turbine Blades and Vanes” discloses prior art methods for weld repairs of air cooled turbine blade tips including squealer tips.
Some gas turbine engine compressor blades are designed so that, during engine operation, the tip portion of the rotating blades rubs a stationary seal or casing, and limits the leakage of working medium gases in the axial flow direction. While the seals are usually more abradable than are the blade tips (so that during such rub interactions, a groove is cut into the seal), the blade tips do wear, and the blades become shorter. As the blades accumulate service time, the total tip wear increases to the point that eventually, the efficiency of the blade and seal system is reduced and cracks may appear in the blades especially at the blade tips such that the blades need to be repaired or replaced. Repairing is much cheaper and more desirable.
The leading and trailing edges and tips of worn blades can be repaired and the airfoils restored to original dimensions by mechanically removing, such as by cutting out or grinding down, the worn and/or damaged areas along the leading and trailing edges and tip of the damaged airfoil and then adding weld filler metal to the tip to build up the leading and trailing edges and tip to a desired dimension using any of several well known welding techniques (typically arc welding techniques) known to those skilled in the art. When an engine is overhauled, compressor blades are either replaced by new parts, which is very expensive, or repaired, which is clearly more desirable if a cost savings may be achieved. Several methods have been devised in which a metal overlay is deposited by spraying or welding metal metallic filler in successive beads onto a substrate to form or dimensionally restore gas turbine engine compressor blade airfoils and, more particularly, the blade's leading and trailing edges and tip. A key limitation to weld repairs is that the repaired parts have a derated life from OEM specs.
Repairing and restoring leading and trailing edges and tip of airfoils by welding causes the airfoil to have a high cycle fatigue HCF capability that is much less than the original equipment manufacturing (OEM) or new part capability. The amount of airfoil that can be repaired and restored by this method is limited because welding causes reduced high cycle fatigue HCF capability. It is highly desirable to repair or restore the leading and trailing edges and tip of airfoils by welding and yet still have a high cycle fatigue HCF capability that as good or nearly as good as that of the original or new part. It is highly desirable to repair or restore a greater amount of the airfoil by welding and yet still have a high cycle fatigue HCF capability that as good or nearly as good as that of the original or new part.
A method of repairing a gas turbine engine airfoil having a periphery that includes leading and trailing edges and a radially outer tip includes machining away airfoil material along at least a portion of the periphery to form at least one cut-back area in the airfoil along at least a portion of at least one of the edges and/or the radially outer tip of the airfoil. Then forming a weldment in the cut-back area by welding successive beads of welding material into the cut-back area beginning with a first bead on a welding surface of the airfoil along the cut-back area and then machining away some of the weld bead material in the weldment to obtain desired finished dimensions of at least one of the edges and/or the radially outer tip of the airfoil. Then imparting deep compressive residual stresses in a pre-stressed region extending into and encompassing the weldment and a portion of the airfoil adjacent the weldment.
An exemplary embodiment of the method further includes machining away airfoil material along only radially outermost portions of the leading and/or trailing edges extending from the outer tip towards a base of the airfoil. This embodiment of the method further includes forming a rounded corner having a semi-circular corner, with an arc and radius of curvature, between the leading edge and/or trailing edge cut-backs and unmachined portions of the airfoil between the outermost portions of the leading and/or trailing edges and the base of the airfoil. A more particular embodiment of the method includes the outermost portions of the leading and/or trailing edges having a spanwise length up to and including about 90% of a span of the airfoil from the outer tip towards the base of the airfoil. The cut-back area may have a maximum cut-back depth up to about 0.22 inches.
The machining away some of the weld bead material in the weldment to obtain desired finished dimensions of at least one of the leading and trailing edges and the radially outer tip of the airfoil may include rough machining and then final finishing of the weldment and the imparting of the deep compressive residual stresses may be performed after the rough machining or after the final finishing of the weldment.
Another exemplary embodiment of the method further includes laser shock peening to impart the deep compressive residual stresses in a pre-stressed region extending into and encompassing the weldment. This exemplary embodiment of the method includes laser shock peening pressure and suction sides of the airfoil and the portion of the airfoil adjacent the weldment.
Another more particular embodiment of the method includes setting a repaired life of a component containing the repaired gas turbine engine airfoil to substantially at or exceeding a new OEM life of the component.
A repaired gas turbine engine airfoil includes the periphery including leading and trailing edges and a radially outer tip, at least one cut-back area in at least a portion of the periphery, the cut-back area being along at least a portion of at least one of the edges and/or the radially outer tip of the airfoil, a weldment including successive beads of welding material in the cut-back area having a first bead on a welding surface of the airfoil along the cut-back area, and deep compressive residual stresses imparted in a pre-stressed region extending into and encompassing the weldment and a portion of the airfoil adjacent the weldment.
A more particular embodiment of the repaired airfoil includes the cut-back area being along at least one of the leading or trailing edges in a radially outermost portion of the leading and/or trailing edges respectively and extending from the outer tip towards a base of the airfoil. The outermost portion of the leading or trailing edges has a spanwise length up to and including about 90% of a span of the airfoil from the outer tip towards the base of the airfoil. A rounded corner is disposed between the leading edge and/or trailing edge cut-backs and unmachined portions of the airfoil between the outermost portions of the leading and/or trailing edges and the base of the airfoil. The rounded corner may be a semi-circular corner having an arc and radius of curvature. The cut-back area may have a maximum cut-back depth up to about 0.22 inches.
A greater degree of damage and/or wear of the leading and trailing edges and tip of compressor blades may be repaired with the present method instead of more expensive replacement of the blades or prior methods of using weldments without imparting compressive residual stresses. The present repair method including imparting deep compressive residual stresses into and encompassing the weldment and a portion of the airfoil adjacent the weldment provides a comprehensive repair process that can more economically repair and dimensionally restore the edges and tips for far greater damaged airfoils.
The foregoing aspects and other features of the invention are explained in the following description, taken in connection with the accompanying drawings where:
Illustrated in
Referring to
A first exemplary embodiment of the repair method disclosed herein is illustrated in
After the airfoil material 50 is machined away weld beads 70 beginning with a first bead 71 on a welding surface 73 of the airfoil along the cut-back area 80, are welded into the cut-back area 80 forming the weldment 82 therein. Typically, airfoil material 50 is removed along only a radially outer half 28 of the airfoil 34, however, in the repair method presented herein, the removal and the cut-back area 80 may extend downwardly to about 90% of the span S from the airfoil outer tip 38 toward the base 32. Then the weldment 82 is machined to near net shape and then finished to final dimensions and surface smoothness.
After the weldment 82 is machined to near net shape or after the weldment 82 is finished to final dimensions and surface smoothness deep compressive residual stresses are imparted in pre-stressed regions 56 extending into and encompassing the weldment 82 and a portion 26 of the airfoil adjacent the weldment 82. Imparting the deep compressive residual stresses in pre-stressed regions 56 is illustrated in the figures as being performed by laser shock peening as indicated by circular spots 58 in
The imparting of deep compressive residual stresses into the weldment allows an extension of permitted maximum cut-back depth 66 to be increased to about 0.2 inches or in a range of 0.18 to 0.22 inches as compared to previous repair methods that allowed only about 0.08 to 0.12 inches from new part dimensions of the leading and trailing edges. Though not drawn to scale, this is illustrated in
The repair method presented herein is also exemplified for gas turbine engine airfoils 34 with worn and/or damaged leading and trailing edges and tip. The repair method is a comprehensive process for restoring the leading and trailing edges and tip of the blade either individually or in combination. Occasionally, but repeatably, the compressor blade 8 rubs on the compressor casing 17 or shroud causing tip damage 52, including burrs, nicks, and tears, on the airfoil outer tip 38 as illustrated in
The periphery 35 of the airfoil 34 is defined by and includes the leading edge LE, the airfoil outer tip 38, and the trailing edge TE. The process is typically preceded by an inspection of the airfoil 34 to determine repairability. After the blade 8 is found to have met repairability requirements, the blade is cleaned and prepped for repair.
Referring to
After the airfoil material 50 is machined away weld beads 70, beginning with a first bead 71 on a welding surface 73 of the airfoil along the cut-back area 80 of the leading edge, trailing edge, and tip cut-backs 62, 63, 64, are welded into the cut-back area 80 forming a weldment 82 therein as illustrated in
Referring to
Compressor and fan blades repaired in this manner using these conventional welding techniques which include TIG (tungsten inert gas) and microTIG can cause defects in and around the welded areas either in the form of porosities and/or microstructural changes. These defects can reduce material fatigue strength. The leading edge of fan and compressor airfoils have a high level of rotational and dynamic stresses. A high pressure compressor (HPC) airfoil is a component doing work on a fluid and there is a very high level of axial stress distributed differentially between the pressure and suction walls of the airfoil. The HPC airfoil, as well as other airfoils in the gas turbine engine, is also subjected to structural damage from solid particles other than the intended fluid flowing across, around and generally into the leading edge of the airfoil. The stress may be due to excitations of the blade in bending and torsional flexure modes. The dominant failure mode may not always be the maximum stress mode but rather a lower stress mode or combination of modes that exist for longer durations over the engine's mission. During engine operation, compressor and fan blades are subject to centrifugal force, aerodynamic force, and vibratory stimuli due to the rotation of the fan and compressor blades over the various operating speeds of the engine. The airfoils of the blades have various modes of resonant vibration (flexure modes) due to the various excitation forces occurring during engine operation. Blades are basically cantilevered from rotor disks and, therefore, may bend or flex generally in the circumferential direction in fundamental and higher order modes of flexure or flex. Airfoils are also subject to fundamental and higher order torsional modes of vibration which occur by twisting around the airfoil span axis. The flex and torsion modes of vibration may also be coupled together further decreasing the life of the blades. To counter these effects on repaired airfoils, the repair method disclosed herein laser shock peens the pressure and suction sides 46, 48 of the airfoil 34 to form laser shock peened patches 86 over the weldment 82 on both the pressure and suction sides 46, 48 of the airfoil 34 either after the near net shape machining step or after the finishing step of after the weldment 82 is welded in. The laser shock peened patches 86 should extend beyond/over the weldment 82 on both the pressure and suction sides 46, 48 of the airfoil 34 as illustrated in
In the exemplary embodiment of the disclosed repair method, the airfoil material 50 along only radially outermost portions 85 of the leading and trailing edges LE, TE extending from the outer tip 38 towards the base of the airfoil is machined away. In previous repair methods, airfoil material along only a radially outer half 28 of the airfoil 34 is machined away, but in the present method with laser shock peening of the weldment, the leading edge and trailing edge cut-backs 62, 63 may extend up to about 90% of the span along the leading and trailing edges.
As further illustrated in
The weldment 82 is machined away to obtain the desired finished dimensions of the leading and trailing edges and radially outer tip by rough and then final blending or finishing of the weldment 82. During the rough machining, the weldment 82 is machined to near net shape and then finished to final dimensions and surface smoothness. Desired finished dimensions of the airfoil's leading edge LE and the airfoil outer tip 38, particularly along the weldment 82, is obtained by contouring of the leading edge LE. Welding parameters and cut-back depths are controlled to prevent airfoil deformation that would require further cold processing to qualify the airfoil for use. The weld beads may be applied with an automated plasma-arc weld process along the cut-back leading and trailing edges and radially outer tip. A Liburdi Laws 500 welding center is one suitable apparatus for the process.
The weldment 82 is subject to loss of high cycle fatigue capability and, thus, the present method includes laser shock peening (LSP) the pressure and suction sides 46, 48 of the airfoil 34 in areas A that entirely encompass the weldment 82. The laser shock peened patches 86 include laser shock peened surfaces 54 formed in the areas A and pre-stressed region 56 having deep compressive residual stresses imparted by laser shock peening (LSP) extending into the airfoil 34 from the laser shock peened surfaces 54. The pre-stressed regions 56 extend beyond the weldment 82 and the leading edge cut-back 62 into the airfoil 34. The laser shock peening may be performed either after the rough or near net machining of the welding material 72 to obtain the near net shape or after final blending or surface finishing to restore the final dimensions of the leading edge LE and the radially outer tip 38. The entire laser shock peened surface 54 is formed by overlapping laser shocked peened circular spots 58.
The laser shock peening induces deep compressive residual stresses in compressive pre-stressed regions 56. The compressive residual stresses are generally about 50-150 KPSI (Kilo Pounds per Square Inch) extending from the laser shocked peened surfaces 54 to a depth of about 20-50 mils into laser shock induced pre-stressed regions 56. The deep compressive residual stresses may also be induced by other cold working methods such as burnishing.
The laser beam shock induced deep compressive residual stresses are produced by repetitively firing a high energy laser beam that is focused on a surface which is covered with paint to create peak power densities having an order of magnitude of a gigawatt/cm.sup.2. The laser beam may be fired through a curtain of flowing water over the laser shock peened surface 54 which is usually painted or otherwise covered with an ablative material and the ablative material is ablated generating plasma which results in shock waves on the surface of the material. These shock waves are re-directed towards the painted surface by the curtain of flowing water to generate travelling shock waves (pressure waves) in the material below the painted surface. The amplitude and quantity of these shockwave determine the depth and intensity of compressive stresses. The ablative material is used to protect the target surface and also to generate plasma but uncoated surfaces may also be laser shock peened. Ablated material is washed out by the curtain of flowing water.
Laser shock peening the weldment in a repaired airfoil as disclosed herein can physically make the airfoil “as good as new”. A key limitation to more conventional weld repairs is that the repaired parts have a derated life from original equipment manufacturer (OEM) specifications. The laser shock peening of the repair weldment as disclosed herein appears to improve and completely overcome the weld debit of the rated life of the repaired component with the weldment in the repaired airfoil. The laser shock peening of the repair weldment may be applied to an airfoil that was originally laser shock peened along the leading and/or trailing edges and/or tip.
While there have been described herein what are considered to be preferred and exemplary embodiments of the present invention, other modifications of the invention shall be apparent to those skilled in the art from the teachings herein and, it is therefore, desired to be secured in the appended claims all such modifications as fall within the true spirit and scope of the invention. Accordingly, what is desired to be secured by Letters Patent of the United States is the invention as defined and differentiated in the following claims.
