US3574924A - Solid state repair method and means - Google Patents

Solid state repair method and means Download PDF

Info

Publication number
US3574924A
US3574924A US3574924DA US3574924A US 3574924 A US3574924 A US 3574924A US 3574924D A US3574924D A US 3574924DA US 3574924 A US3574924 A US 3574924A
Authority
US
Grant status
Grant
Patent type
Prior art keywords
articles
dies
masses
article
mass
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Expired - Lifetime
Application number
Inventor
Gordon L Dibble
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Boeing Co
Original Assignee
Boeing Co
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Grant date

Links

Images

Classifications

    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01DNON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/005Repairing methods or devices
    • BPERFORMING OPERATIONS; TRANSPORTING
    • B22CASTING; POWDER METALLURGY
    • B22DCASTING OF METALS; CASTING OF OTHER SUBSTANCES BY THE SAME PROCESSES OR DEVICES
    • B22D19/00Casting in, on, or around objects which form part of the product
    • B22D19/10Repairing defective or damaged objects by metal casting procedures
    • YGENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02TCLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00Aeronautics or air transport
    • Y02T50/60Efficient propulsion technologies
    • Y02T50/67Relevant aircraft propulsion technologies
    • Y02T50/671Measures to reduce the propulsor weight
    • YGENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02TCLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00Aeronautics or air transport
    • Y02T50/60Efficient propulsion technologies
    • Y02T50/67Relevant aircraft propulsion technologies
    • Y02T50/673Improving the blades aerodynamics
    • YGENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y10TECHNICAL SUBJECTS COVERED BY FORMER USPC
    • Y10TTECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
    • Y10T29/00Metal working
    • Y10T29/49Method of mechanical manufacture
    • Y10T29/49718Repairing
    • Y10T29/49732Repairing by attaching repair preform, e.g., remaking, restoring, or patching
    • Y10T29/49734Repairing by attaching repair preform, e.g., remaking, restoring, or patching and removing damaged material
    • Y10T29/49737Metallurgically attaching preform

Abstract

A method of rebuilding a jet engine compressor, fan, or turbine blade containing a worn or damaged area by trimming off the portion of such blade surrounding the area and replacing such portion with another of precisely corresponding size and metallurgical composition. Standardized templates define the size of the trimmed off portion. The replaced portion is integrally joined to the blade by solid state molecular diffusion bonding in a mold exactly duplicating the contour of the original blade when new.

Description

United States Patent Gordon L. Dibble Fontana, Calif.

Oct. 28, 1968 Apr. 13, 1971 North American Rockwell Corporation lnventor Appl. No. Filed Patented Assignee SOLID STATE REPAIR METHOD AND MEANS 9 Claims, 8 Drawing Figs.

U.S. Cl 29/401, 29/475, 22/498 Int. Cl B22d 19/10, B23p 7/00 Field of Search 29/ 401 475, 156.8 (B),493,498,402, 198, 504; 52/514 References Cited UNITED STATES PATENTS 276,499 4/1883 Story 144/310X 1,397,167 11/1921 52/51'4 2,203,389 6/1940 83/565X 3,058,202 10/1962 29/1 56.8 3,256,598 6/1966 29/493X 3,281,923 1l/1966 29/475X 3,335,488 8/1967 29/493X 3,487,530 l/1970 29/402 Primary Examiner-John F. Campbell Assistant Examiner-Victor A. DiPalma Attorneys-Wil1iam R. Lane, Charles F. Dischler and Harold H. Card, Jr.

ABSTRACT: A method of rebuilding a jet engine compressor, fan, or turbine blade containing a worn or damaged area by trimming off the portion of such blade surrounding the area and replacing such portion with another of precisely corresponding size and metallurgical composition. Standardized templates define the size of the trimmed off portion. The replaced portion is integrally joined to the blade by solid state molecular difiusion bonding in a mold exactly duplicating the contour of the original blade when new.

Patented April 13, 1971 2 Sheets-Sheet 1 INVENTOR.

BY 6 1mm 0/55 A Trozavzr Patented April 13, 1911 3,514,924

2 Sheets-Sheet 2 INVENTOR. 60200 1.. 0/554;

xm k

- A rive/var SOLID STATE REPAIR METHOD AND MEANS BACKGROUND OF THE INVENTION Engines of turbojet or turbofan-type such as used in modern military and commercial aircraft contain many compressor and turbine blades, each of which is precision formed. Engine performance and reliability depend directly upon the detailed design contour and structural integrity of these blades. In axial flow compressors of some jet engines, as many as 20 stages of compression are accomplished by a corresponding number of blade rows, each containing from 50 to 150 individual blades. The moving rows are separated by stationary alternate rows in close proximity, whereby dimensional accuracy of the blades is crucially important. Similarly, in turbofan engines, each individual blade is contoured for maximum efficiency at the design rotation speed, whereby the angle of attack at each radial location along the blade length decreases as the tangential velocity increases. Accuracy of such contour is crucially important in achieving proper aerodynamic performance of the blade. In addition, structural integrity of the blades is absolutely essential due to the high centrifugal stresses and elevated temperatures to which they are exposed during high speed rotation under normal operating conditions.

When a blade becomes damaged due to entrance of a foreign object, erosion, temperature or load stress, such as to produce local distortion or microscopic cracking, it is unsafe for aircraft use. Accordingly, compressor and turbofan blades which are found to contain any abnormal deviation or defect in shape, dimension, structural integrity, or surface smoothness are completely discarded and replaced by new blades. Due to the extreme care and expense involved in fabrication of such blades and the costly materials used therein, replacement of such blades is the outstanding major item of expense in jet and turbofan engine maintenance.

SUMMARY OF THE INVENTION The invention in this case provides a method for repairing metallic articles which are found to contain localized defects such as cracks, pits, dents, surface erosion, worn areas, or other dimensional deficiencies from any cause. As illustratively applied to turbofan blades, the novel method begins with a machining step to cut away that portion of the blade surrounding the defect. The foregoing step is carefully controlled to produce a cutout portion of predetermined size and shape according to a set of standardized template patterns. For economy and mass-production convenience, a large number of incremental elements of standard size and shape conforming to the mentioned templates are prefabricated for use in rebuilding damaged blades. After the mentioned machining or cutting step, the blade is placed into a two-piece mold having workfaces which define a cavity oppositely corresponding to the precise contour of a complete undamaged blade. One or more of the mentioned prefabricated incremental elements is positioned in the cutout area of the damaged blade. Heat and pressure are applied to the two-piece mold as required to effect gradual deformation of the incremental element into conformance with the stated workface contour, and to unite the incremental mass to the adjoining blade material integrally by a solid-state molecular diffusion bond.

DESCRIPTION OF DRAWINGS FIG. 5 shows a finished rebuilt blade following completion of the novel process;

FIG. 6 shows a general perspective view of a retort useful in the bonding step of the novel process; and

FIGS. 7 and 8 are cross-sectional views taken along line 7-7 of FIGS. 4 and 6 showing the tooling and workpiece elements in two stages of the bonding operation used to produce the rebuilt blade of FIG. 5.

DETAILED DESCRIPTION Referring to the drawings described above, use of the novel process disclosed herein may be seen in connection with repairing or rebuilding turbofan blades such as blade 10 in FIG. 1. In a typical case, blade 10 is provided with a strong base portion 12 with flanges or similar means to key the same securely in a rotating hub or the like (not shown). Blade 10 is an airfoil with a progressively changing angle of attack between base I2 and distal end or tip 14 in the familiar manner of propeller blades. The leading and trailing edges of blade I0 are designated 16 and 18, respectively. Intermediate blade portions I2 and 14, blade 10 is provided with stabilizing means in the form of platform projections extending in opposite directions from both sides of the blade, only one of which is visible in FIG. I and denoted by reference numeral 20. When many blades identical to blade 10 are mounted in a complete circle around a hub, the mentioned intermediate platforms 20 contact each other between each pair of adjoining blades to form a circular segmented flange, the segments of which comprise platforms 20. Thus, each of the platforms is designed to bear firmly against the next adjacent platform and thereby stabilize the same against shock or vibration effects in the individual blades during high speed rotation of the turbofan.

In turbofans of conventional type, blade 10 may have an overall length on the order of l2 to 15 inches or more. The airfoil contours of blade I0 are precision engineered and fabricated to produce maximum efficiency of engine performance at the particular speed and pressure conditions at which the engine incorporating such blades is designed to operate. Any dents, punctures or other localized deformations in the contours of blade I0 seriously compromise the aerodynamic properties of the blade and are completely intolerable. In the illustrative case of FIG. 1, blade 10 contains damaged portion 22 such as frequently experienced in aircraft due to entrance of a foreign object in the inlet airflow of a turbofan engine during operation thereof. Under present commercial airline engine overhaul procedures, such damage would require blade 10 to be discarded and replaced by a new blade in the absence of the inventive process disclosed herein.

Referring to FIG. 2, the process fcr repairing blade 10 begins by cutting or other suitable operations whereby the portion of blade 10 containing damaged area 22 is removed from the blade. The portion thus removed has a predetermined size and shape conforming to a template 24. In theillustrative case of blade 30 shown in FIG. 3, erosion damage along substantially the entire length of leading edge 16 such as may be encountered in a typical case is shown. Blade 30 may be seen to contain other damaged portions 32 and 34 in addition to leading edge erosion area 36. Any change in normal dimensions, such as due to wear of a blade from any cause during service use thereof may be overcome by the repair method disclosed herein. In addition to template 24, it may be seen particularly from FIG. 4 that additional templates 25, 26, 27, and 28 are adapted to define those portions of blade 30 requiring removal from the blade in order to insure that no damaged areas remain on blade 30 after the cutting operation is completed. The portions of blade 30 to be thus removed are determined by placing templates 2428 in the locations identified by reference numerals 44-48, respectively, and scribing or otherwise marking blade 30 to trace the outline of the templates thereon.

It will be understood by those skilled in the art that templates 24-28 represent standardized areas of predetermined size'and location which define those portions of blades and 30 which most frequently sustain damage during service use of the same, and that the templates are adapted for repair of blades 10 and 30 on a mass production basis. Thus, those portions of blades 10 and 30 which are removed during the stated cutting operations will be replaced by incremental masses of the same material used in the mentioned blades, such masses having essentially the same shape and surface area as templates 2428, but not necessarily the same thickness. Viewing FIG. 4, for example, the incremental mass required to replace the material removed from blade 30 and defined by area 46 will have the same shape and location as area 46, and is designated by reference numeral 60 in FIG. 7.

Referring to FIGS. 7 and 8, the novel method in this case involves use of restraining die or mold means such as illustratively shown by mating dies 50 and 52. Dies 50 and 52 may be fabricated by any suitable process known to the prior art, such as by casting from plaster patterns or cutting operations on metallic blocks. Dies 50 and 52 when in completely mated relationship enclose a cavity 54 thcrebetween defined by the workfaces 56 and 58 of dies 50 and 52, respectively. Cavity 54 has a contour oppositely and precisely corresponding to the surfaces of a complete new blade identical to blade 10 or 30 but without any damage or defects. Following the cutting operations described in connection with FIG. 4, for example, blade 30 is positioned in lower die 52 with incremental masses corresponding in size and location to areas 44-48 shown in FIG. 4 preplaced in the manner thus suggested, Incremental mass 60, for example, seen in FIG. 7 occupies the area of the void produced by removal of material defined by area 46 in FIG. 4. However, while mass 60 has substantially the same peripheral dimensions as area 46, the thickness 1 of mass 60 is uniform throughout the mass and is of particular significance in practicing the method disclosed herein.

In FIG. 7, blade 30 with the damaged portions removed therefrom is positioned between mating dies 50 and 52 within a surrounding retort 62. Mass 60 is positioned within cavity 54 in the void resulting from removal of area 46 from blade 30, other incremental masses being similarly preplaced in the same pattern suggested by area 4448 in FIG. 4. Each of the stated incremental masses has substantially uniform thickness, as suggested by thickness t of mass 60, but not necessarily the same thickness as between different masses. Thus, the thickness of mass 60, for example, is predetermined to result in a total volumetric content of the mass which will coincide almost exactly with the volume of the void within cavity 54 resulting from removal of area 46 of blade 30 when the blade is positioned between dies 50 and 52, and the dies are fully mated together in the relationship shown by FIG. 8. Ideally, the volumetric content of mass 60 should equal that of the mentioned void, In no case should mass 60 ever be sized to result in less volume than such void. Therefore, any tolerance allowed for dimensional inconsistencies in the manufacture of masses 60 must be limited to oversizing thereof rather than undersizing, with respect to the mentioned ideal volumetric content. As a practical matter, very slight oversizing of mass 60, as well as all the other incremental masses discussed above, on the order of 0.01 percent to 10 percent in excess of the stated ideal volumetric content, is preferable in order to compensate for any inconsistencies in the cutting operations necessary to prepare blade 30 for repair. It is, of course, the purpose of templates 24-28 to standardize the cutting patterns and minimize such inconsistencies, but some slight variations during lengthy mass production runs will inevitably occur.

With the workpiece and tooling elements arranged within retort 62 as shown by FIG, 7 and discusses above, a vacuum source (not shown) is preferably connected to the retort through conduit 64 provided for this purpose The retort and its contents are heated to a suitable temperature for difi'usion bonding to occur between the material in blade 30 and the incremental masses adjacent thereto such as mass 60. The

amount of such heating will depend upon the amount of pressure to which retort 62 and its contents will be subjected and the duration of its exposure to the conditions identified with solid-state molecular diffusion bonding of the workpiece materials. Of the various materials suitable for use in turbofan blades, titanium or an alloy thereof is widely used because of its high strength and light weight, although other and different materials may be used in various other types of article repaired by the novel process disclosed herein. Where the workpiece material is titanium, complete bonding together with the necessary creep deformation required to reshape mass 60 into confon'nity with the contours of die cavity 54, may be accomplished at l600 F. and 1000 psi. compressive force continuously maintained for 24 hours in the direction suggested by arrows 66 and 68. in FIG. 7, for example.

It will be understood that the inventive concept in this case may be practiced with a wide variation of metals and alloys and with articles of different size and shape, and that the parameters for achieving solid-state diffusion bonding will necessarily vary for each particular choice of workpiece material. Among the metals or alloys which may be joined by solid-state diffusion bonding are aluminum, stainless steel, titanium, nickel, tantalum, molybdenum, zirconium, and columbium. Diffusion bonding is characterized by intermolecular exchange between contacting surfaces of the workpiece at suitable pressures and at temperatures below the melting point of the workpiece material. In some cases, a thin interleaf material, or eutectic former, is provided while in other forms of solid-state bonding no interleaf material is necessary. The prior art involving solid-state or intermolecular diffusion bonding includes issued US. Pat. Nos. 3,145,466; 3, l 80,022; 3,044,160; 2,850,798; and 3,170,234. The precise values of time-temperature and pressure utilized in connection with bonding workpiece materials is not a critical or limiting feature of the broad concept disclosed herein, but specific materials and parameters are stated for illustration only. Similarly, many different metals or alloys for dies 50 and 52 could be used to practice the inventive principles taught herein, although hard tool steel such as those high in nickel and cobalt content are preferable due to the relatively severe temperatures and load conditions to which the dies are subjected. Illustratively, 4l30 steel widely used commercially is suitable for the dies.

When deformation of mass 60 is complete, dies 50 and 52 will be in the completely mated position suggested by FIG. 8 wherein cavity 54 is fully occupied by the rebuilt blade 70. In the foregoing relationship, a slight gap between the dies may be seen to exist as indicated by reference numeral 72. Compressive force denoted by arrows 66 and 68 is transmitted through the dies and directly to the workpiece elements contained within cavity 54. It is this compressive force which produces deformation of incremental masses such as mass 60 as required to confonn the same into the final desired shape of the rebuilt blade 70. Moreover, application of compressive force vertically through mass 60, for example, will be understood to cause lateral or horizontal force to be exerted by mass 60 against the contacting portions of blade 30 and die cavity 54. Thus, in the absence of lateral restraint offered by blade 30 and cavity 54, the compressive force applied to mass 60 would produce considerable increase in the length and width thereof simultaneous with reduction in its thickness t. Since lateral deformation is restrained, a considerable reaction force transversely through mass 60 occurs due to vertical pressure applied to retort 62. It is this lateral component of force which is essential for solid-state molecular diffusion bonding to occur between the contacting surfaces of blade 30 and mass 60. If the volumetric content of mass 60 were insufficient to occupy completely the void produced by removal of material from blade 30, complete deformation of mass 60 would occur without creation of the necessary lateral force. It is the foregoing consideration which makes the dimensions of the incremental masses such as mass 60 so critical in practicing the process disclosed herein.

Following completion of the bonding steps discussed above, retort 62 is opened and dies 50 and 52 are separated from each other to permit removal of blade 70. It will be understood that blade 70 is a single solid unitary mass of material completely homogenous throughout with regard to metallurgical properties and composition. The shape of blade 70 oppositely corresponds in every detail with the contours of cavity 54 when dies 50 and 52 are fully mated as seen in FIG. 8. The foregoing feature completely avoids any necessity for machining operations on blade 70, resulting in a very significant savings of time, money, and material, the importance of which increases with use of costly materials such as titanium. 1n practicing the process thus disclosed, manual deburring of edges is occasionally required to finish rebuilt blades, but every such blade is identical to all others formed within the same set of matched dies 50 and 52.

In further connection with the illustrative case of turbofan blade repair using the method disclosed herein, use of the alloy known as Ti-6Al-4V in such blades is common. This alloy has the following approximate composition by weight:

Aluminum 5% to 6% percent Vanadium 3 5a to 4% percent Carbon 0.08 to 0.1 percent Hydrogen 0.010 to 0.012 percent Titanium balance When the foregoing material is used in both blade 30 and mass 60 as well as all other incremental masses required to occupy cavity 54 completely in FIG. 7, for example, retort 62 and its contents are heated to a temperature of from 1650 to 1725 F. with a vacuum of 1X10 millimeters of Mercury or less maintained in the retort. The stated temperature is continued for a period of from about 8 to about 16 hours, while a pressure of from about 2000 to about 5000 psi. is applied in the direction of arrows 66 and 68 for the stated period. Where only slight deformation of mass 60 and the other incremental masses is involved, such as by preshaping the same to the approximate aerodynamic curvature of the final part before the bonding step, the lower end of the forementioned pressure and time period ranges are applicable.

lclaim: 1. A method of repairing a worn or damaged metallic article, comprising the steps of:

providing a pair of mating dies, the workfaces of which are adapted to define a cavity having the precise size and shape of said article in the new and undamaged condition when said dies are in mating relationship; removing a portion of said article to create a vacant area thereon, said portion including the worn or damaged area of said article; placing said article in said dies with at least one separate incremental mass positioned in said. vacant area, said mass being of the same material as said article, the volumetric content of said mass being at least equal to the volume of the void fonned by said vacant area in said cavity when said dies are in mating relationship; and applying sufficient heat and pressure to said dies to cause said mass both to deform into conformity with said void and to join said article by solid-state molecular diffusion bonding therebetween. 2. The method set forth in claim 1 above, further including, prior to said removing step:

providing at least one template for-mass production use on a plurality of damaged articles corresponding generally with said damaged article; placing said template in close proximity to said damaged article over said damaged area; and tracing said template on said article to define said vacant area whereby said removed portion corresponds in plan view with said template. 3. The method set forth in claim 1 above wherein: said mass has a volumetric content from 0.01 to 10.0

percent larger than said volume of said void. -4.- Ihemet Set qrthjn hin! labors where n;

said metallic article is made of Ti-6Al-4V titanium alloy;

and

said heat and pressure are from about 1600 to 1725 from 2000 to 5000 p.s.i., respectively.

5. The method set forth in claim 4 above, wherein:

said dies containing said article and said mass are positioned within an airtight retort; and

a vacuum is maintained in said retort continuously during application of said heat and pressure.

6. A mass-production method for repairing a multitude of worn or damaged metallic articles having identical design but a variety of different types and locations of wear or damage, comprising the steps of:

forming a plurality of standardized templates having predetermined area dimensions;

prefabricating a plurality of incremental masses corresponding in area with said area dimensions of said templates;

removing portions of said articles, said removed portions containing said worn or damaged areas and corresponding in area with said templates;

placing said masses in the voids resulting in said articles from removal of said portions; and

joining said masses to said articles by solid-state molecular diffusion bonding therebetween.

7. The method set forth in claim 6 above, wherein:

said bonding is accomplished by the steps of placing said individual articles together with said masses between a pair of mating dies defining a cavity having the contour of said articles in the undamaged condition;

heating said dies and said articles to a temperature sufficient for said bonding to occur at a coordinated pressure; and applying said pressure to said dies to compress said articles therebetween until said bonding occurs.

8. The method set forth in claim 6 above, wherein:

said masses have a total volumetric content from 0.01 to 10.0 percent larger than the total volumetric content of said removed portions W V I l fii m e th od of mass production rebuilding a plurality of worn or damaged metallic articles of identical design,

comprising the steps of:

forming a plurality of standardized templates having predetermined plan area dimensions;

prefabricating a plurality of incremental corresponding in plan view with said templates;

trimming said articles to remove portions therefrom, said removed portions containing said worn or damaged locations and corresponding in plan area with said templates;

providing a pair of mating dies, the workfaces of which are adapted to define a cavity having the precise size and shape of said articles in the new or undamaged condition when said dies are in fully mated relationship;

successively placing each of said trimmed articles between said dies with a plurality of said masses positioned in the vacant areas resulting in said articles from removal of said portions therefrom, said masses being of the samematerial as said articles, the volumetric content of said masses in combination being at least equal to the volume of the voids formed in said cavity due to removal of said portions when said dies contain said trimmed articles and are in fully mated relationship;

surrounding said dies with an airtight retort;

maintaining a vacuum within said retort;

heating said retort and its contents to a temperature sufficient for solid-state molecular diffusion bonding between said incremental masses and said trimmed articles;

applying sufficient pressure to said dies to compress said articles and masses therebetween in an amount sufficient to deform said masses into conformity with said cavity and to achieve said diffusion bonding; and

masses said masses having an initial combined volumetficonteh t" from 0.01 to 10.0 percent larger than said volume of said voids in said cavity resulting from said removal of said portions.

Claims (9)

1. A method of repairing a worn or damaged metallic article, comprising the steps of: providing a pair of mating dies, the workfaces of which are adapted to define a cavity having the precise size and shape of said article in the new and undamaged condition when said dies are in mating relationship; removing a portion of said article to create a vacant area thereon, said portion including the worn or damaged area of said article; placing said article in said dies with at least one separate incremental mass positioned in said vacant area, said mass being of the same material as said article, the volumetric content of said mass being at least equal to the volume of the void formed by said vacant area in said cavity when said dies are in mating relationship; and applying sufficient heat and pressure to said dies to cause said mass both to deform into conformity with said void and to join said article by solid-state molecular diffusion bonding therebetween.
2. The method set forth in claim 1 above, further including, prior to said removing step: providing at least one template for mass production use on a plurality of damaged articles corresponding generally with said damaged article; placing said template in close proximity to said damaged article over said damaged area; and tracing said template on said article to define said vacant area whereby said removed portion corresponds in plan view with said template.
3. The method set forth in claim 1 above wherein: said mass has a volumetric content from 0.01 to 10.0 percent larger than said volume of said void.
4. The method set forth in claim 1 above wherein: said metallic article is made of Ti-6A1-4V titanium alloy; and said heat and pressure are from about 1600* to 1725* F. and from 2000 to 5000 p.s.i., respectively.
5. The method set forth in claim 4 above, wherein: said dies containing said article and said mass are positioned within an airtight retort; and a vacuum is maintained in said retort continuously during application of said heat and pressure.
6. A mass-production method for repairing a multitude of worn or damaged metallic articles having identical design but a variety of different types and locations of wear or damage, comprising the steps of: forming a plurality of standardized templates having predetermined area dimensions; prefabricating a plurality of incremental masses corresponding in area with said area dimensions of said templates; removing portions of said articles, said removed portions containing said worn or damaged areas and corresponding in area with said templates; placing said masses in the voids resulting in said articles from removal of said portions; and joining said masses to said articles by solid-state molecular diffusion bonding therebetween.
7. The method set forth in claim 6 above, wherein: said bonding is accomplished by the steps of placing said individual articles together with said masses between a pair of mating dies defining a cavity having the contour of said articles in the undamaged condition; heating said dies and said articles to a temperature sufficient for said bonding to occur at a coordinated pressure; and applying said pressure to said dies to compress said articles therebetween until said bonding occurs.
8. The method set forth in claim 6 above, wherein: said masses have a total volumetric content from 0.01 to 10.0 percent larger than the total volumetric content of said removed portions.
9. The method of mass production rebuilding a plurality of worn or damaged metallic articles of identical design, comprising the steps of: forming a plurality of standardized templates having predetermined plan area dimensions; prefabricating a plurality of incremental masses corresponding in plan view with said templates; trimming said articles to remove portions therefrom, said removed portions containing said worn or damaged locations and corresponding in plan area with said templates; providing a pair of mating dies, the workfaces of which are adapted to define a cavity having the precise size and shape of said articles in the new or undamaged condition when said dies are in fully mated relationship; successively placing each of said trimmed articles between said dies with a plurality of said masses positioned in the vacant areas resulting in said articles from removal of said portions therefrom, said masses being of the same material as said articles, the volumetric content of said masses in combination being at least equal to the volume of the voids formed in said cavity due to removal of said portions when said dies contain said trimmed articles and are in fully mated relationship; surrounding said dies with an airtight retort; maintaining a vacuum within said retort; heating said retort and its contents to a temperature sufficient for solid-state molecular diffusion bonding between said incremental masses and said trimmed articles; applying sufficient pressure to said dies to compress said articles and masses therebetween in an amount sufficient to deform said masses into conformity with said cavity and to achieve said diffusion bonding; and said masses having an initial combined volumetric content from 0.01 to 10.0 percent larger than said volume of said voids in said cavity resulting from said removal of said portions.
US3574924A 1968-10-28 1968-10-28 Solid state repair method and means Expired - Lifetime US3574924A (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
US77122068 true 1968-10-28 1968-10-28

Publications (1)

Publication Number Publication Date
US3574924A true US3574924A (en) 1971-04-13

Family

ID=25091105

Family Applications (1)

Application Number Title Priority Date Filing Date
US3574924A Expired - Lifetime US3574924A (en) 1968-10-28 1968-10-28 Solid state repair method and means

Country Status (1)

Country Link
US (1) US3574924A (en)

Cited By (30)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4008844A (en) * 1975-01-06 1977-02-22 United Technologies Corporation Method of repairing surface defects using metallic filler material
US4051585A (en) * 1976-07-26 1977-10-04 United Technologies Corporation Method of forming a turbine rotor
US4121894A (en) * 1975-09-15 1978-10-24 Cretella Salvatore Refurbished turbine components, such as vanes or blades
US4161056A (en) * 1977-08-05 1979-07-17 P.R.K., Inc. Method and device for repairing damaged screw propellers
US4214355A (en) * 1977-12-21 1980-07-29 General Electric Company Method for repairing a turbomachinery blade tip
US4305697A (en) * 1980-03-19 1981-12-15 General Electric Company Method and replacement member for repairing a gas turbine engine vane assembly
DE3109342A1 (en) * 1980-03-19 1981-12-24 Gen Electric A method of repairing a gas turbine engine blade luftgekuehlten and provided for this purpose ersatzteil
EP0056328A2 (en) * 1981-01-12 1982-07-21 Refurbished Turbine Components Limited Turbine blade repair
US4611744A (en) * 1982-06-23 1986-09-16 Refurbished Turbine Components Ltd. Turbine blade repair
US4614296A (en) * 1981-08-26 1986-09-30 Societe Nationale D'etude Et De Construction De Moteurs D'aviation S.N.E.C.M.A. Diffusion brazing process for pieces of superalloys
US4874290A (en) * 1988-08-26 1989-10-17 Solar Turbines Incorporated Turbine blade top clearance control system
EP0401187A2 (en) * 1989-06-01 1990-12-05 Abb Stal Ab Method for reconstruction of blades and vanes in steam turbines at existing erosion damages
US5242102A (en) * 1992-12-14 1993-09-07 Nicolas Raymond G Method for forming and diffusion bonding titanium alloys in a contaminant-free liquid retort
WO1996000840A1 (en) * 1994-06-30 1996-01-11 United Technologies Corporation Turbine vane flow area restoration method
US5755031A (en) * 1996-11-12 1998-05-26 United Technologies Corporation Method for attaching a rotor blade to an integrally bladed rotor
WO1999039097A1 (en) * 1998-01-30 1999-08-05 Voith Hydro Gmbh & Co. Kg Method for producing a wear-endangered component of a turbo-machine
US6339878B1 (en) 2000-03-27 2002-01-22 United Technologies Corporation Method of repairing an airfoil
US6394750B1 (en) * 2000-04-03 2002-05-28 United Technologies Corporation Method and detail for processing a stator vane
US20040172825A1 (en) * 2003-03-03 2004-09-09 Memmen Robert L. Turbine element repair
US20040172827A1 (en) * 2003-03-03 2004-09-09 Kinstler Monika D. Fan and compressor blade dovetail restoration process
US20070163115A1 (en) * 2006-01-16 2007-07-19 United Technologies Corporation Turbine platform repair using laser clad
US20080057254A1 (en) * 2003-03-03 2008-03-06 United Technologies Corporation Turbine element repair
US20080213092A1 (en) * 2007-03-01 2008-09-04 Honeywell International, Inc. Repaired vane assemblies and methods of repairing vane assemblies
US20080236536A1 (en) * 2007-03-30 2008-10-02 Caterpillar Inc. Cast engine component having metallurgically bonded inserts
US20090320287A1 (en) * 2005-12-15 2009-12-31 United Technologies Corporation Compressor blade flow form technique for repair
US20100162565A1 (en) * 2008-12-30 2010-07-01 Mukherji Tapas K Refurbishing method and system for a main rotor blade spar
US8956700B2 (en) 2011-10-19 2015-02-17 General Electric Company Method for adhering a coating to a substrate structure
CN105881888A (en) * 2016-04-07 2016-08-24 山东英特力新材料有限公司 Method for mounting equipment on composite material sandwich panel
CN105881931A (en) * 2016-04-07 2016-08-24 山东英特力新材料有限公司 Complementation method for hole of large sandwich-structure composite cabin
CN105904747A (en) * 2016-04-13 2016-08-31 山东英特力新材料有限公司 Repairing method of large composite material shelter

Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US276499A (en) * 1883-04-24 Ephraim j
US1397167A (en) * 1920-01-19 1921-11-15 Hopper Charles Inlay for repair of scored engine-cylinders and method of making same
US2203389A (en) * 1939-11-13 1940-06-04 Bausch & Lomb Apparatus for cutting contours
US3058202A (en) * 1958-06-27 1962-10-16 Edward A Stalker Method of making hollow blades for compressors, turbines, and the like
US3256598A (en) * 1963-07-25 1966-06-21 Martin Marietta Corp Diffusion bonding
US3281923A (en) * 1964-08-27 1966-11-01 Corning Glass Works Method of attaching leads to thin films
US3335488A (en) * 1963-09-20 1967-08-15 Gen Dynamics Corp Material joining by solid state cohesion
US3487530A (en) * 1967-10-09 1970-01-06 Abex Corp Method of repairing casting defects

Patent Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US276499A (en) * 1883-04-24 Ephraim j
US1397167A (en) * 1920-01-19 1921-11-15 Hopper Charles Inlay for repair of scored engine-cylinders and method of making same
US2203389A (en) * 1939-11-13 1940-06-04 Bausch & Lomb Apparatus for cutting contours
US3058202A (en) * 1958-06-27 1962-10-16 Edward A Stalker Method of making hollow blades for compressors, turbines, and the like
US3256598A (en) * 1963-07-25 1966-06-21 Martin Marietta Corp Diffusion bonding
US3335488A (en) * 1963-09-20 1967-08-15 Gen Dynamics Corp Material joining by solid state cohesion
US3281923A (en) * 1964-08-27 1966-11-01 Corning Glass Works Method of attaching leads to thin films
US3487530A (en) * 1967-10-09 1970-01-06 Abex Corp Method of repairing casting defects

Cited By (41)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4008844A (en) * 1975-01-06 1977-02-22 United Technologies Corporation Method of repairing surface defects using metallic filler material
US4121894A (en) * 1975-09-15 1978-10-24 Cretella Salvatore Refurbished turbine components, such as vanes or blades
US4051585A (en) * 1976-07-26 1977-10-04 United Technologies Corporation Method of forming a turbine rotor
US4161056A (en) * 1977-08-05 1979-07-17 P.R.K., Inc. Method and device for repairing damaged screw propellers
US4214355A (en) * 1977-12-21 1980-07-29 General Electric Company Method for repairing a turbomachinery blade tip
US4305697A (en) * 1980-03-19 1981-12-15 General Electric Company Method and replacement member for repairing a gas turbine engine vane assembly
DE3109342A1 (en) * 1980-03-19 1981-12-24 Gen Electric A method of repairing a gas turbine engine blade luftgekuehlten and provided for this purpose ersatzteil
DE3110180A1 (en) * 1980-03-19 1982-02-18 Gen Electric "Procedure to repair a luftgekuehlten gas turbine engine vane assembly, and in return provided casting spare parts"
US4326833A (en) * 1980-03-19 1982-04-27 General Electric Company Method and replacement member for repairing a gas turbine engine blade member
EP0056328A2 (en) * 1981-01-12 1982-07-21 Refurbished Turbine Components Limited Turbine blade repair
EP0056328A3 (en) * 1981-01-12 1984-03-28 Refurbished Turbine Components Limited Turbine blade repair
US4614296A (en) * 1981-08-26 1986-09-30 Societe Nationale D'etude Et De Construction De Moteurs D'aviation S.N.E.C.M.A. Diffusion brazing process for pieces of superalloys
US4611744A (en) * 1982-06-23 1986-09-16 Refurbished Turbine Components Ltd. Turbine blade repair
US4874290A (en) * 1988-08-26 1989-10-17 Solar Turbines Incorporated Turbine blade top clearance control system
EP0401187A2 (en) * 1989-06-01 1990-12-05 Abb Stal Ab Method for reconstruction of blades and vanes in steam turbines at existing erosion damages
EP0401187A3 (en) * 1989-06-01 1992-10-07 Abb Stal Ab Method for reconstruction of blades and vanes in steam turbines at existing erosion damages
US5242102A (en) * 1992-12-14 1993-09-07 Nicolas Raymond G Method for forming and diffusion bonding titanium alloys in a contaminant-free liquid retort
WO1996000840A1 (en) * 1994-06-30 1996-01-11 United Technologies Corporation Turbine vane flow area restoration method
US5755031A (en) * 1996-11-12 1998-05-26 United Technologies Corporation Method for attaching a rotor blade to an integrally bladed rotor
WO1999039097A1 (en) * 1998-01-30 1999-08-05 Voith Hydro Gmbh & Co. Kg Method for producing a wear-endangered component of a turbo-machine
US6339878B1 (en) 2000-03-27 2002-01-22 United Technologies Corporation Method of repairing an airfoil
US6394750B1 (en) * 2000-04-03 2002-05-28 United Technologies Corporation Method and detail for processing a stator vane
US20040172825A1 (en) * 2003-03-03 2004-09-09 Memmen Robert L. Turbine element repair
US20040172827A1 (en) * 2003-03-03 2004-09-09 Kinstler Monika D. Fan and compressor blade dovetail restoration process
US7216428B2 (en) * 2003-03-03 2007-05-15 United Technologies Corporation Method for turbine element repairing
US8122600B2 (en) * 2003-03-03 2012-02-28 United Technologies Corporation Fan and compressor blade dovetail restoration process
US20080057254A1 (en) * 2003-03-03 2008-03-06 United Technologies Corporation Turbine element repair
US20100196684A1 (en) * 2003-03-03 2010-08-05 United Technologies Corporation Turbine Element Repair
US20090320287A1 (en) * 2005-12-15 2009-12-31 United Technologies Corporation Compressor blade flow form technique for repair
US8127442B2 (en) * 2005-12-15 2012-03-06 United Technologies Corporation Compressor blade flow form technique for repair
US20070163115A1 (en) * 2006-01-16 2007-07-19 United Technologies Corporation Turbine platform repair using laser clad
US20080213092A1 (en) * 2007-03-01 2008-09-04 Honeywell International, Inc. Repaired vane assemblies and methods of repairing vane assemblies
US7959409B2 (en) * 2007-03-01 2011-06-14 Honeywell International Inc. Repaired vane assemblies and methods of repairing vane assemblies
US20080236536A1 (en) * 2007-03-30 2008-10-02 Caterpillar Inc. Cast engine component having metallurgically bonded inserts
US20100162565A1 (en) * 2008-12-30 2010-07-01 Mukherji Tapas K Refurbishing method and system for a main rotor blade spar
US8800145B2 (en) * 2008-12-30 2014-08-12 Sikorsky Aircraft Corporation Refurbishing method and system for a main rotor blade spar
US8956700B2 (en) 2011-10-19 2015-02-17 General Electric Company Method for adhering a coating to a substrate structure
CN105881888A (en) * 2016-04-07 2016-08-24 山东英特力新材料有限公司 Method for mounting equipment on composite material sandwich panel
CN105881931A (en) * 2016-04-07 2016-08-24 山东英特力新材料有限公司 Complementation method for hole of large sandwich-structure composite cabin
CN105881931B (en) * 2016-04-07 2018-03-16 山东英特力新材料有限公司 Pods composite sandwich structure A large aperture completion method
CN105904747A (en) * 2016-04-13 2016-08-31 山东英特力新材料有限公司 Repairing method of large composite material shelter

Similar Documents

Publication Publication Date Title
US5666643A (en) High temperature braze material
US5622638A (en) Method for forming an environmentally resistant blade tip
US2687278A (en) Article with passages
US6905728B1 (en) Cold gas-dynamic spray repair on gas turbine engine components
US5141400A (en) Wide chord fan blade
US4050133A (en) Method of refurbishing turbine vanes and the like
US6568077B1 (en) Blisk weld repair
US3940268A (en) Method for producing rotor discs
US4817858A (en) Method of manufacturing a workpiece of any given cross-sectional dimensions from an oxide-dispersion-hardened nickel-based superalloy with directional coarse columnar crystals
US4051585A (en) Method of forming a turbine rotor
US20070084047A1 (en) Method for repairing and/or modifying component parts of a gas turbine
US5142778A (en) Gas turbine engine component repair
US5111570A (en) Forge joining repair technique
US6416278B1 (en) Turbine nozzle segment and method of repairing same
US3736638A (en) Method for bonding opposed parts of a hollow article together
US5038014A (en) Fabrication of components by layered deposition
US5100050A (en) Method of manufacturing dual alloy turbine disks
US4214355A (en) Method for repairing a turbomachinery blade tip
US6467168B2 (en) Method of manufacturing an article by diffusion bonding and superplastic forming
US5867885A (en) IBR fixture and method of machining
US7841834B1 (en) Method and leading edge replacement insert for repairing a turbine engine blade
US6394750B1 (en) Method and detail for processing a stator vane
US20050220995A1 (en) Cold gas-dynamic spraying of wear resistant alloys on turbine blades
US4882823A (en) Superplastic forming diffusion bonding process
US5113583A (en) Integrally bladed rotor fabrication