HK1231013B - Method of repairing and manufacturing of turbine engine components and turbine engine components - Google Patents

Description

Method for repairing and manufacturing turbine engine component and turbine engine component

Technical Field

The present invention relates to fusion welding and can be used to repair and fabricate turbine engine components made of nickel, cobalt and iron-based superalloys using manual and automated welding methods such as gas tungsten arc welding (GTAW), laser welding (LBW), electron beam welding (EBW), plasma welding (PAW) and micro plasma welding (MPW).

Background Art

The present invention relates to fusion welding and can be used to repair and manufacture various turbine engine components, more particularly turbine blades made of equiaxed polycrystalline, single crystal and directionally solidified superalloys using cladding and fusion welding processes.

In fusion welding, the coalescence or connection between two or more products is carried out by melting the base material with or without the introduction of fillers, and then cooling and crystallizing the weld pool. Fusion welding can produce properties equivalent to those of the base material under a wide range of temperatures and conditions. However, solidification and the retention of residual stresses often lead to cracks, making it difficult to weld Inconel713, Inconel738, Rene77, Rene80, Rene142, CMSX-4, ReneN4, ReneN5 and other high γ' superalloys with low ductility and prone to cracking in the liquefaction heat affected zone (HAZ).

Brazing can produce crack-free joints because it does not require melting of the base material to achieve coalescence. Brazing is performed by melting and solidifying only the brazing material. However, at high temperatures, the mechanical properties of the brazed joint are typically 50-75% lower than those of the base material.

The poor mechanical properties of brazed joints formed from most nickel and cobalt braze materials are related to the high boron content in these materials and do not allow for large-scale restoration of turbine blades and structural repair of other engine components.

Therefore, despite its tendency to crack, welding is often used more frequently than brazing to manufacture and repair various products, including turbine engine components. However, according to US Pat. No. 5,897,801, to avoid cracking during fusion welding, turbine blades made of materials with low ductility are preheated to temperatures exceeding 900°C prior to welding. Welding is accomplished by striking an arc in a preselected area to locally melt the parent material, providing a filler metal having the same composition as the nickel-based superalloy of the product, and feeding the filler metal into the arc. The arc causes the filler metal to melt and fuse with the parent material, forming a weld deposit upon solidification.

A similar approach is also used in the method disclosed in US 6659332. The article is repaired by removing the damaged material present in the defect area and then preheating the article to 60-98% of the solidus temperature of the base material in a chamber containing a shielding gas and then welding.

In order to minimize the weld stresses in the blade due to the large amount of heat energy applied during fusion welding, the blade is subjected to controlled heating before the weld repair and controlled cooling after the weld repair according to the method described in CA1207137.

Preheating of turbine blades increases repair costs and does not guarantee crack-free welding due to the low ductility of components made from precipitation-hardened superalloys.

Therefore, currently only preheating to temperatures exceeding 900 °C allows crack-free welding on precipitation-hardened equiaxed polycrystalline and directionally solidified high γ′ phase superalloys.

Therefore, one of the main objects of the present invention is to develop a new cost-effective method for repairing engine components by welding and cladding on polycrystalline, directionally solidified and single crystal superalloys at ambient temperature.

Summary of the Invention

We have found that a preferred embodiment of the method of repairing and manufacturing turbine engine components of the present invention includes: a pre-weld preparation step of removing damaged material and contaminants to expose a defect-free base material; welding the damaged area to repair the damaged area by preferably two dissimilar fillers using a fusion welding process selected from laser, micro-beam plasma, plasma, electron beam and gas tungsten arc welding, wherein the first dissimilar filler is selected from ductile nickel-based and cobalt-based alloys, including high-temperature dendrites and low-temperature interdendritic eutectics, and having a solidus temperature below the solidus temperature of the base material due to a 0.05 wt.% to 1.2 wt.% boron additive; and then, after applying the transition layer, the base material is welded to the weld metal at a temperature above the aging temperature of the base material. temperature) but below the initial melting temperature of the base material, and performing a diffusion heat treatment for about 30 minutes to about 24 hours; applying a top anti-oxidation layer using a fusion welding process and a second dissimilar filler, wherein the second dissimilar filler comprises about 5 to 12 wt.% Co, about 12 to 25 wt.% Cr, about trace to 5 wt.% Mo, about trace to 5 wt.% W, about 1 wt.% to 5 wt.% Ti, about trace to 0.1 wt.% Zr, about trace to about 1.5 wt.% Hf, about trace to 0.2 wt.% B, about 3 to 6 wt.% .% Al, about 0.5 wt.% to about 6 wt.% Si, about trace to about 5.5 wt.% Re, about trace to about 4 wt.% Ta, and nickel, and the balance impurities; restoring the original geometry of the engine component by a method selected from machining, grinding, and polishing after a post-weld heat treatment of the base material selected from hot isostatic pressing, annealing, aging, and stress relieving; non-destructive testing; and dimensional inspection, as well as other embodiments of the present invention discussed below; producing defect-free welds and HAZs on various high gamma-' precipitation-hardened nickel-base superalloys at ambient temperature.

According to another embodiment, the filler used to apply the transition layer to a base material containing about trace amounts to about 3.5 wt. % aluminum is selected from a nickel-based alloy containing about 0.05 wt. % to about 0.6 wt. % boron.

According to another embodiment, the filler used to apply the transition layer to a base material containing about 3 wt.% to about 8.0 wt.% aluminum is selected from a nickel-based alloy containing about 0.4 wt.% to about 1.2 wt.% boron.

Another preferred embodiment includes the additional step of machining the transition layer to a uniform thickness of 0.3 mm or more.

According to another preferred embodiment, in order to restore the base material and improve the mechanical properties at the weld, the repaired engine component is subjected to hot isostatic pressing treatment in any of the following cases, depending on the state of the base material, before welding, after applying the transition layer, or after applying the top oxidation-resistant layer.

According to a further preferred embodiment, in order to improve weldability and perform vacuum cleaning, the turbine engine component is subjected to an annealing heat treatment in vacuum or in a protective gas, preferably hydrogen, before the application of the transition layer.

In order to simplify the use of automated welding and cladding and to allow the top oxidation resistant layer to be applied directly to the repair area, defective material at the tip of the blade is removed by cutting off the tip at least 0.25 mm below a typical repairable damage to the turbine blade, thereby allowing the top oxidation resistant layer to be applied directly during the subsequent repair of the previously applied transition layer.

According to a preferred embodiment aimed at improving the properties of the base and weld materials, the turbine blade is subjected to a post-weld heat treatment selected from annealing, precipitation hardening of the base material or both or stress relieving using parameters selected from those prescribed for the base material.

According to a preferred embodiment of the present invention, the fusion welding process is performed at ambient temperature.

However, if necessary, based on weldability and scrap rate statistics of the base material according to another embodiment, the fusion welding process may be performed by preheating to a temperature of about 600°C to about 1100°C.

Preferred embodiments of the present invention may be used to repair and fabricate turbine engine components made from single crystal, directionally solidified, equiaxed nickel, cobalt, and iron-based superalloys.

All preferred embodiments can be used to repair and manufacture turbine engine parts selected from nozzle guide vanes (NGV), compressor stator blades, compressor blades, high-pressure compressor (HPC) blades, high-pressure turbine (HPT) blades, low-pressure turbine (LPT) blades, shroud rings, seal segments, casings, guide plates, combustion chambers, flame tubes, fuel nozzles, and manifolds of aviation and industrial turbine engines.

The present invention is based on a further improvement in the repair of turbine engine components using welding materials described in pending patent applications WO2015095949, CA2850698 (CN104511702(A)), PCT/CA2014/000752 and WO2014063222 of Liburdi Engineering Ltd. The combination of a boron-containing ductile welding material for the transition layer, followed by a diffusion heat treatment and the application of a silicon-containing top oxidation-resistant layer is a key step in producing crack-free components with excellent oxidation resistance and distinguishing the present invention from the prior art.

According to the present invention, the following advantages can be observed:

The method has been found to enable crack-free welding at ambient temperature on most polycrystalline, directionally solidified and single crystal superalloys with a high content of γ′ phase, which reduces costs, increases productivity and improves the health and safety of working conditions.

Due to the addition of silicon and the optimization of the Al-Si-Cr ratio, the repaired engine components exhibit excellent oxidation resistance that exceeds that of most base materials.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG1 is a micrograph of the fusion zone of a test specimen made of IN738 with a transition layer formed by GTAW-MA welding using a first dissimilar filler alloy, 3669-6B, showing HAZ liquid crack healing via low interdendritic eutectics during weld pool solidification.

FIG2 shows the microstructure of the crack-free heat-affected zone of the trailing edge of an IGT blade made of GTD111DS superalloy, the transition layer adjacent to the base material produced by GTAW-MA welding using a first dissimilar filler alloy 3687B, and the top oxidation-resistant layer produced using a second dissimilar filler alloy 3667S.

FIG3 shows the defect-free microstructure of the top oxidation-resistant layer produced using the second dissimilar filler alloy 3667S at a magnification of x500.

Figure 4 is a micrograph of a crack-free HAZ and transition layer produced using GTAW-MA and the first dissimilar filler alloy, 3669-6B.

Standard acronyms

AMS – Space Material Specification (Standard)

ASTM-American Society for Testing and Materials (standards)

AWS-American Welding Society (Standard)

OEM - Original Equipment Manufacturer

NDT – Non-destructive testing

PWHT-Post Weld Heat Treatment

HAZ-Heat Affected Zone

IGT-Industrial Turbine Engine

LPT-Low Pressure Turbine

HPT-High Pressure Turbine

HPC-High Pressure Compressor

NGV – Nozzle Guide Vanes

GTAW – Gas Tungsten Arc Welding

PAW-Plasma Arc Welding

MPW-Micro Plasma Welding

LBW-Laser Beam Welding/Laser Welding

EBW - Electron Beam Welding

HIP-Hot Isostatic Pressing

EDM-Electrical Discharge Machining

EM-Engine Manual

SPM-Standard Practical Manual

UTS-Ultimate Tensile Strength

DS - Directional Solidification (alloy or material)

TE – Nozzle guide vanes and blade trailing edges

LE – Nozzle guide vanes and leading edges

DTA-Differential Thermal Analysis

EDS - Energy Dispersive X-ray Spectroscopy

EPMA - Electron Probe Microanalysis

Vocabulary and terms (definitions)

Alloy - A metallic compound consisting of a mixture of two or more materials. Superalloy - A metallic material that has oxidation resistance and mechanical properties at high temperatures.

Nickel-based superalloys - materials in which the nickel content exceeds the content of other alloying elements.

Wrought nickel alloy - a nickel-based alloy that has been bent, hammered, forged, or physically deformed into a desired shape. Wrought nickel alloys are often welded under the same conditions as certain types of steel.

Cast Nickel Alloy – An alloy containing nickel that is poured or cast in liquid form into a mold and cooled into a solid shape.

Base Metal or Material – One of two or more metals to be welded together to form a joint.

Cracking – A break that forms in a weld during or after the weld pool has solidified.

Ductility – The ability of metals and alloys to be stretched, elongated, or formed without breaking.

Hardness - The ability of metals and alloys to resist indentation, penetration, and scratching.

Heat treatment – a controlled heating and cooling process used to alter the structure of a material and change its physical and mechanical properties.

Solution Heat Treatment – A heat treatment method in which an alloy is heated to a specific temperature and held for a period of time to dissolve one or more alloying elements into solid solution, followed by rapid cooling.

Ageing or hardening - Hardening caused by precipitation of a component from a supersaturated solid solution.

Aging or precipitation hardening heat treatment - artificial aging in which a component/constituent precipitates from a supersaturated solid solution due to heating and exposure to high temperatures.

Multi-step aging heat treatment – is a process in which the heat treatment temperature is gradually reduced during the heat treatment to achieve the desired precipitation morphology and superalloy properties.

First Aging - The first high temperature stage of a multi-stage aging heat treatment.

Secondary Aging - The second stage of a multi-step aging heat treatment, conducted at a lower temperature than the primary aging temperature for the selected superalloy.

Hot Isostatic Pressing (HIP) - is the simultaneous application of high temperature and high pressure to metals and other materials for a specific amount of time to improve their mechanical properties.

Overaging – Aging at times and temperatures greater than those required to obtain the maximum change in some property, so that the property is altered in the direction of the initial value, which is particularly applicable to altering the properties of turbine engine components exposed to high temperatures under operating conditions of turbine engine components made of precipitation hardened superalloys.

Diffusion heat treatment - heating to diffuse a component, especially boron, in a solid, especially a base material, with the intention of making the composition of the weld and base material uniform across all parts, especially engine components.

Argon Quenching – Argon gas is introduced into a vacuum heat treatment chamber at the annealing temperature, causing the alloy to cool rapidly to ambient temperature.

Weldability – The ability of a material to be welded into a specified, suitable configuration under the conditions imposed and to perform well for its intended purpose.

Non-weldable Material - Material that cannot be welded using a fusion welding process at ambient temperature.

Welding powder – welding material in powder form that is added during the process of forming a weld joint or cladding weld.

Welding Wire – welding material in the form of a wire that is added during the formation of a weld joint or cladding weld.

Welding Rod – Welding wire cut to standard lengths.

Cladding – The process of applying a relatively thick (>0.5 mm (0.02 in.)) layer of weld material and/or composite weld powder with minimal penetration into the base material to improve wear and/or corrosion resistance or other properties and/or to restore a component to desired dimensions.

Weld – A localized joining of metals or nonmetals produced by heating the materials to the welding temperature with or without the application of pressure, or by the application of pressure alone with or without the use of welding material.

Weld Bead – A weld produced in one pass.

Heat Affected Zone (HAZ) - The portion of the base metal that is not melted, but whose mechanical properties or microstructure are altered by the heat of welding.

Dilution – A change in the chemical composition of the weld material produced by mixing base metal or previous weld metal in the weld bead, measured as a percentage of base metal or previous weld metal in the weld bead.

Welding – The process of joining materials used to create a weld.

Fusion Welding - A welding process that uses the melting of the base metal to form the weld.

Gas tungsten arc welding (GTAW) - According to the AWS definition, it is an arc welding process that produces a metallic bond by heating the metal with an electric arc between a tungsten (non-consumable) electrode and the workpiece (also called the base material). Shielding is provided by a gas or gas mixture. Pressure may or may not be applied, and filler metal may or may not be used.

Plasma Arc Welding (PAW) - According to the AWS definition, it is an arc welding process in which the metals are joined by heating the metals by a converging arc between the electrode and the workpiece (base metal) (also called a transferred arc) or by an arc between the electrode and a converging nozzle (also called a non-transferred arc).

Laser welding and cladding (LBW) - according to the AWS definition, it is a welding process that uses heat obtained by applying a concentrated coherent beam to bombard the joint or base material, respectively, to produce a material bond.

Weld Pass – A single progression of a welding or cladding operation along a joint, weld deposit, or substrate. The result of a weld pass is a weld bead, weld layer, or spray deposit.

Multi-pass Cladding and Welding – A weld/joint formed by more than two passes.

Weld Defect – A discontinuity or interruption resulting from natural or cumulative effects that causes a part or product to fail to meet the minimum applicable acceptance criteria or specifications.

Discontinuity – An interruption in the typical structure of the weld metal, such as a lack of uniformity in the mechanical, metallurgical, or physical properties of the base or weld metal.

Linear discontinuity – a weld defect with an aspect ratio of 3:1 or greater.

Crack – A fracture-type discontinuity characterized by a sharp tip and a high aspect ratio, usually exceeding three.

Solidification Shrinkage – The volumetric shrinkage of a metal during solidification.

Crack - A small crack-like discontinuity in the fracture surface with only a slight separation (opening displacement). The prefix macro- or micro- indicates relative size.

Weld pool – the local volume of molten metal in a weld before it solidifies.

Carbides – compounds composed of carbon and less electronegative elements. Carbon can form carbides with metals (such as chromium, niobium, molybdenum, tantalum, titanium, tungsten, and other metals from Groups IVB, VB, and VIB) and non-metals (such as boron, calcium, or silicon). Metal carbides are characterized by their extreme hardness and resistance to high temperatures.

Boride – A compound composed of two elements, with boron being the more electronegative element. Boron forms borides with metals and nonmetals.

Gamma Phase – The continuous matrix (called gamma) is a face-centered cubic (fcc) nickel-based austenitic phase that typically contains high proportions of solid solution elements such as Co, Cr, Mo, and W.

Austenite – a solid solution of more than one element in a face-centered cubic phase.

γ' phase – the primary strengthening phase in nickel-base superalloys, is a compound consisting of nickel and aluminum or titanium (Ni3Al or Ni3Ti) precipitated congruently in the austenitic γ matrix.

Ultimate Tensile Strength (UTS) - A material's resistance to longitudinal stress, measured as the minimum amount of longitudinal stress required to cause the material to break.

Yield Strength - The ability of a metal to withstand a progressive force without permanent deformation.

Creep - is the tendency of a solid material to slowly move or permanently deform under the influence of stress. Creep occurs when metals are subjected to continuous tensile loads at high temperatures.

The fracture test is conducted according to ASTM E139 by applying a continuous load to a tensile specimen held at a constant temperature. Fracture tests are conducted in a similar manner to creep tests, but at higher stress levels. The specimen is tested until failure, and the time to failure is measured. The time until fracture at the specified load characterizes the fracture properties of the material.

Fracture Strength – The nominal stress at which a material breaks, which is not necessarily equal to the ultimate strength.

Recrystallization – is the formation of a new, strain-free grain structure from an existing grain structure, usually accompanied by grain growth during heating.

Recrystallization Temperature – is the approximate temperature at which complete recrystallization of the existing grain structure occurs within a specified time.

Crack-free weld—a weld bead that contains no linear markings having a length to width ratio equal to or greater than 3:1, as detected by nondestructive testing or metallographic examination of the weld at magnifications up to 100°.

Differential Thermal Analysis (DTA) is a thermal analysis technique, similar to differential scanning calorimetry, in which a sample under investigation and an inert reference sample are subjected to identical thermal cycles, while any temperature difference between the sample and the reference is recorded. This differential temperature is then plotted against time or temperature (a DTA curve, or thermogram). Changes in the sample, whether exothermic or endothermic, can be detected relative to the inert reference.

DTA curve - is a curve that will provide data about the transformations that occur in a sample such as melting, solidification, phase changes, and sublimation. The area under the DTA peak is the enthalpy change and is not affected by the heat capacity of the sample.

Energy Dispersive X-ray Spectroscopy (EDS) - is an analytical technique used for elemental analysis or chemical characterization of samples.

Electron Microprobe Analyzer (EMPA) - is an analytical tool used to non-destructively determine the chemical composition of small volumes of solid materials.

DETAILED DESCRIPTION

HPT and LPT blades for aerospace and IGT engines, as well as other turbine engine components, are manufactured from superalloys, directionally solidified materials, and single-crystal materials with low ductility to ensure high creep and durability performance. However, low ductility increases the susceptibility of these materials to cracking, hindering the accommodation of residual stresses through plastic deformation. In addition to the aforementioned, most equiaxed and directionally solidified superalloys, including the most common GTD111 and IN738, are susceptible to liquefaction cracking in the HAZ along the fusion line.

The method of the present invention solves the cracking problem of the base material while enhancing the oxidation resistance of the repaired turbine blade by applying a transition layer of a first dissimilar boron-containing ductile filler. The solidus temperature of the filler is lowered to below that of the base material by boron, thereby preventing overheating of the base material and enabling self-healing of cracks through redistribution of interdendritic eutectics by capillary forces. Diffusion heat treatment performed after application of the transition layer causes boron to diffuse into the base material along the fusion line, thereby increasing the base material's crack resistance. Application of a top oxidation-resistant layer utilizes a second dissimilar filler with a silicon additive and an optimized Al-Si-Cr ratio, thereby ensuring excellent oxidation resistance of the repaired component. A turbine engine component heat treatment method selected from HIP, annealing, and aging, or a combination of all of the above, and PWHT restores the base material's original properties.

The method of the present invention is disclosed by way of an example of repairing a turbine blade made of GTD111 equiaxed superalloy, and then by way of an example of repairing actual engine components made of equiaxed and directionally solidified (DS) GTD111 and IN738 superalloys. These materials were chosen for illustration due to their high sensitivity to overheating, their susceptibility to HAZ liquefaction cracking, and their widespread use in the manufacture of HPT blades and NGVs for IGT engines.

Prior to welding repair, the turbine blades are stripped and cleaned of protective coatings/platings according to relevant standard practices.

After cleaning, the turbine blades undergo fluorescent penetrant inspection (FPI) and dimensional inspection according to AMS2647, followed by tip grinding to remove defective material and expose defect-free base material. For automated welding, defective material is uniformly removed by grinding to the maximum crack depth or at least 0.25 mm below the typical crack depth, enabling direct welding repair using a second dissimilar filler in a subsequent repair process. For manual welding, it is acceptable to remove a single crack by removing the skin.

The condition of the base material is assessed through stress-rupture testing, which aims to optimize the scope and sequence of repair work. At least one blade in each group is cut into machined stress-rupture specimens according to ASTM E-8. Stress-rupture testing of these specimens is performed according to ASTM E-139.

If the base material's properties—particularly ductility—have been severely deteriorated due to creep and microvoid formation in engine components exposed to high stresses and temperatures, HIP is performed prior to welding. Turbine blades are heated in an inert gas, typically argon, which applies isostatic pressure uniformly in all directions. This causes the material to become "plastic," allowing voids to collapse under differential pressure. The void-diffused surfaces bond together to effectively eliminate defects, achieving near-theoretical density while improving the blade's mechanical properties. The parameters for HIP processing are typically specified in various manuals and relevant OEM specifications, as well as numerous publications. For example, according to "OM Study of Effect of HIP and Heat Treatments on Microstructural" by Panyawat Wangyao, Viyaporn Krongtong, Weerasak Homkrajai et al., in the journal Metals, Materials and Minerals, Vol. 17, No. 1, pp. 87-92, 2007, HIP of GTD111 superalloy was performed at a temperature higher than the annealing temperature for 4 hours.

Annealing heat treatment of turbine blades made of GTD111 superalloy is performed at 1000°C for 1 hour to facilitate vacuum cleaning after HIP. Annealing after HIP can restore the ductility of the base material to a level at which it can withstand the thermal stresses generated by welding using a first dissimilar filler material with a reduced solidus temperature and high ductility without cracking.

At least the transition layer is applied by a fusion welding process selected from the group consisting of laser, microplasma, plasma, electron beam, and gas tungsten arc welding, and a first dissimilar filler material. The first dissimilar filler material, according to a preferred embodiment, comprises approximately 10 to 25 wt.% Cr, trace amounts of approximately 10 wt.% Co, trace amounts of approximately 1.5 wt.% Al, trace amounts of approximately 20 wt.% Fe, trace amounts of approximately 1 wt.% Si, trace amounts of approximately 0.2 wt.% C, trace amounts of approximately 3.5 wt.% Ti, approximately 0.05 wt.% to approximately 1.2 wt.% B, a total of approximately 2 wt.% to 25 wt.% of at least one element selected from the group consisting of niobium, molybdenum, and tungsten, nickel, and the remainder as impurities. The boron content of the first filler material depends on the aluminum content of the base material. The GTD111 alloy contains 3.5 wt.% aluminum. As determined experimentally, in order to eliminate liquefaction cracking in the HAZ of GTD111 and produce welds with sufficient ductility by manual GTAW welding to allow for accommodation of residual stresses by plastic deformation of the weld metal, the first dissimilar filler should contain about 0.4 wt.% to 1.2 wt.% B and specifically 21.5 wt.% Cr-9 wt.% Mo-3.7 wt.% Nb-0.5 wt.% B and nickel and the balance of impurities (as described in PCT/CA2014/000752), which is further designated as alloy 3698-6B. Another prominent first dissimilar filler in the form of welding wire for manual and automatic GTAW and LBW is nickel-based alloy 3687B, the chemical composition of which is described in WO2014063222, including 0.5 wt.% B, 0.2 wt.% C and other alloying elements.

To improve the weldability of the base material, provide stress relief, and minimize boron diffusion from the transition layer into the top oxidation-resistant layer, after applying the transition layer, the component undergoes a diffusion heat treatment, or a combination of diffusion and primary aging heat treatment, at a temperature above the base material's aging heat treatment temperature but below its initial melting point. Experimental results indicate that sufficient boron diffusion into the base material to a depth of 0.3-1 mm can be observed at a temperature of 1205-1220°C for two hours. Blades made from GTD111 superalloy (similar to IN738, as described in Matthew J. Donachie and Stephen J. Donachie, "Superalloys," A Technical Guide, 2nd Edition, ASM International, 2002, p. 141) were then subjected to primary aging. After the diffusion and primary aging heat treatments, the blade tip was machined to achieve near-uniform thickness of the transition layer, followed by fluorescent penetration (FPI) of the weld and HAZ. The first filler material contained 0.5 wt.% B. Boron is a melting point depressant and causes the formation of a composite structure comprising high temperature interconnected scaffolds and low temperature interdendritic eutectics as the weld pool solidifies.

According to DTA, the solidus temperature of Alloy 3669-6B is 1201.45°C, which is much lower than the solidus temperature of IN738 of about 1284°C, while the solidus temperature of the high-temperature dendrites containing about 0.03 wt% B is 1295.4°C. According to EDS analysis, the boron content in the interdendritic eutectic and borides is about 0.9 wt.% in the welded condition and reaches as high as 9.5 wt.% after diffusion cycling and first aging.

The solidus temperature of the dendrites exceeds even that of the base material, making welds formed using the first dissimilar filler extremely stable at high temperatures. The abundant low-temperature eutectics during weld pool solidification allow melt microcracks to self-heal along the fusion line via capillary action, as shown in Figure 1. Consequently, despite weld metal formation at the high solidus temperature of the dendrites, both the weld metal and the HAZ of the base material remain crack-free due to weld pool solidification. Furthermore, as demonstrated in Example 1, the material of the transition layer exhibits excellent ductility, which improves its adaptability to solidification and thermal stresses generated by plastic deformation within the weld metal.

The first aging does not reduce the ductility of the transition layer because the content of γ' phase forming elements such as aluminum and titanium in the first dissimilar filler is negligible. The first aging is carried out for two hours, followed by quenching to ambient temperature with argon.

After FPI, turbine blades are undergoing intensive degreasing, visual and dimensional inspection as per relevant standard specifications.

For applying the top oxidation-resistant layer, automated LBW or MPW is used in combination with a second, dissimilar filler in powder form. GTAW-MA manual welding is more technically and cost-effective with wire. Since high-strength γ' superalloys are difficult to extrude, a combination of filler in powder form and laser or microplasma welding is more cost-effective.

It has been found that in order to maintain reasonable costs for a second dissimilar filler alloy 3669-B in the form of welding powder and welding wire, the welding powder and welding wire should contain about 3.5 wt.% Al, with other alloying elements in the amounts of about 14 wt.% Cr, 10 wt.% Co, 3.5 wt.% Mo, 4.5 wt.% W, 0.15 wt.% C, 0.02 wt.% B, and 1.1 wt.% Si, with Ni and the balance being impurities (as described in WO2015095949). The use of this filler alloy at a temperature of 995°C results in an oxidation resistance that is 4-5 times that of the GTD111 alloy shown in Example 1. For automated LBW or MPB or EBW weld repairs, using a second dissimilar nickel-based filler powder containing about 0.01 wt.% B, 1.8 wt.% Si (also known as alloy 3667S) will produce better results. This alloy is a preferred choice for hot and harsh environments, and the filler powder alloy 3653BS contains 0.2 wt.% B, 1.2 wt.% Si and nickel and the balance of impurities for repairing structural parts such as HPT, LPT, NGV components.

After applying a top oxidation-resistant layer using a second dissimilar filler selected based on the operating conditions, the weldability of the base material, and the requirements for the weld's mechanical properties and oxidation resistance, the turbine engine component undergoes a secondary aging heat treatment using parameters selected from those specified for the base material. The secondary aging heat treatment of the GTD111 superalloy is performed at 845°C for 24 hours in vacuum or under a shielding atmosphere, followed by an argon quench. The aging heat treatment at 845°C does not significantly diffuse boron from the transition layer into the top oxidation-resistant layer, ensuring excellent performance at the blade tip, which is exposed to high temperatures.

After heat treatment, the repair area is machined to the required dimensions using electrical discharge, conventional grinding or hand grinding, followed by polishing and super finishing of the blade surface.

Turbine blade dimensional inspection according to relevant EM standards, as well as FPI according to AMS2647 and radiographic inspection according to ASTME192-04 or relevant repair instructions are used to ensure that all repaired engine components have met the specified requirements.

Example 1

The turbine blades were manufactured from an equiaxed GTD111 superalloy, which is extremely susceptible to liquefaction cracking in the HAZ. The concave and convex sides of the blades were protected by oxidation-resistant coatings and were able to withstand the operating conditions without significant damage to the base material. However, the unprotected tip of the blades exhibited severe oxidation. The turbine blades were repaired in accordance with a preferred embodiment of the present invention using manual GTAW-MA welding with a welding current of approximately 60-80A and an arc voltage of 12-14V using two dissimilar materials. Argon gas was used to protect the weld area.

Specimens for mechanical testing were produced using the same welding parameters, a first exotic ductile filler alloy 3698-6B in wire form containing 0.4 wt. % B, and a second exotic filler alloy 3669-S1 in wire form having improved oxidation resistance through optimization of aluminum, chromium, and silicon contents.

Tensile specimens fabricated from the base material, transition layer, and top oxidation-resistant layer were tested at 20°C and 982°C to evaluate the transition layer's ability to accommodate residual stresses caused by plastic deformation during welding and to resist nucleation and thermal fatigue crack growth under operating conditions. Tensile testing of the specimens was conducted according to ASTM E-8 at room temperature and ASTM E-21 at 982°C.

Cyclic oxidation tests were conducted using samples with a diameter of 5 mm and a length of 25 mm extracted from the weld. The samples were heated to 995°C in air, then held at this temperature for 50 minutes, and then air-cooled to below 400°C for 500 cycles. The sample weight was measured before and after the test.

Evaluation of the mechanical properties confirmed that the base material could withstand the application of the transition layer by standard GTAW-MA welding at ambient temperature using the first dissimilar filler alloy 3698-6B. Therefore, neither HIP nor annealing heat treatments were required.

The diffusion heat treatment was carried out for two hours at the annealing temperature of the base material, followed by quenching with argon and application of an approximately 3 mm thick transition layer followed by initial aging in vacuum.

After the initial aging, the component was subjected to FPI, and the transition layer was machined to a thickness of approximately 1.6 mm and subjected to standard degreasing. An approximately 2.5 mm thick oxidation-resistant layer was applied by two passes of GTAW-MA welding using a second dissimilar filler alloy, 3669-S1, in the form of a 1.14 mm diameter wire, followed by a secondary aging at 845°C in vacuum for 24 hours. Welding of the top oxidation-resistant layer was also completed at ambient temperature.

Weld quality was assessed using NDT, including standard FPI and radiographic testing. A verification specimen (test specimen) and one blade from the set were also subjected to destructive testing by extracting two samples, approximately 18–22 mm long, from the trailing edge and mid-section of the tip for metallographic examination. The verification specimen passed metallographic examination. The weld and HAZ were crack-free and met acceptance criteria.

At the final stage of repair, the turbine blades undergo cutting-edge machining and dimensional inspection. Verification samples undergo tensile testing to prove the repair procedure is satisfactory.

As shown in Table 1 below, the transition layer has high ductility at room temperature and excellent ductility at 982°C, significantly exceeding that of the base material. The top oxide layer has excellent tensile strength and ductility at both 20°C and 982°C, as well as excellent oxidation resistance, exceeding that of the base material at 995°C, as shown in Table 2.

Table 1 Mechanical properties of transition layer and top oxidation resistant layer

Table 2 Oxidation resistance of the base material and the top oxidation-resistant layer, mass loss after 500 cycles at 995 °C.

Material Mass loss, g GTD 111 base material 0.213 Weld metal produced using a second dissimilar filler alloy, 3669-S1 0.058

According to NDT, the welds have met the acceptance criteria.

Testing of the repaired blades under engine conditions showed that the repaired area had excellent resistance to thermal fatigue cracking due to the high ductility of the transition layer and the high oxidation resistance of the top layer – which was enhanced by optimizing the Al-Si-Cr ratio.

Example 2

The turbine blades of the new generation IGT engine are made of a more advanced directionally solidified GTD111DS superalloy, which has better strength in the radial direction and is more resistant to axial cracking. As a preferred embodiment, it allows the use of a first filler alloy 3687B, which contains 0.5wt.% B, 0.2wt% C and other alloying elements in the form of welding wire for applying a transition layer by GTAW.

The base material of this group was affected by long-term exposure to service conditions beyond standard serviceability limits. Therefore, to restore the base material, the blades were HIPed and then vacuum annealed after tip preparation and before welding. The transition layer was applied using a first dissimilar filler alloy, 3687B, in the form of welding wire, GTAW-MA welding, and the welding parameters described in Example 1. After welding, the transition layer was flattened to prepare for the application of a top oxidation-resistant layer using a second dissimilar filler powder by LBW. After the transition layer was applied, the blades were diffused and subjected to a first aging heat treatment to prevent boron diffusion into the top oxidation-resistant layer during the final PWHT process.

To apply the top oxidation-resistant layer, a second dissimilar filler alloy, 3667S, was used in powder form. The top oxidation-resistant layer was applied via three-pass LBW welding on a Liburdi LAWS 500 system at a welding speed of 1.5 mm/s, a laser beam power of 420 W, a powder feed rate of 6.5 g/min, a laser beam oscillation speed of 20 mm/s, and argon shielding. Verification test specimens for mechanical testing were fabricated using the same welding and heat treatment parameters and filler material. The weld repair was performed at ambient temperature.

After application of the top oxidation resistant layer, the blades and check samples were subjected to a secondary aging heat treatment at 845°C for twenty-four (24) hours, and then the blade tips were restored by machining, EDM and grinding.

Final FPI, radiographic, and dimensional inspections were conducted to verify that the repaired blades met relevant specifications. Verification samples were subjected to tensile testing at 20°C and 982°C. Accelerated cyclic oxidation testing was performed by heating 5mm diameter and 25mm long samples extracted from the weld metal to 1120°C, holding at this temperature for 60 minutes, and then air cooling to below 400°C. The oxidation resistance of the samples was comparable to that of turbine blades made from Advanced Aerospace Turbine Engine Material (AATEM), which has superior oxidation resistance to CDT111 DS.

A randomly selected blade was also subjected to destructive testing. A longitudinal sample was taken from the trailing edge. A transverse sample was taken from the leading edge. Both samples were polished to a roughness of 0.5 μm and etched using the standard etchant Marble's. Figure 2 shows the structure of a defect-free weld and base material, and the DS structure of the base material, transition layer, and top oxidation-resistant layer. Figure 3 shows the microstructure of a defect-free top oxidation-resistant layer at a higher magnification.

Table 3 Mechanical properties of transition layer and top oxidation resistant layer

Table 4 Oxidation resistance of AATEM and top oxidation-resistant layer, mass loss after 100 cycles at 1120 °C

Material Mass loss, g AATEM -0.0376 3667S -0.0002

Example 3

A NGV made of difficult-to-weld IN738 was repaired to demonstrate the applicability of the method to the restoration of non-rotating airfoils. Under operating conditions, the NGV developed cracking and significant thinning of the trailing edge (TE) due to erosion and thermal cycling.

Prior to repair, the aluminized coating is stripped from the NGV, followed by chemical cleaning and FPI inspection. Cracks and defective material are removed by descaling the trailing edge, followed by FPI to verify crack removal.

To improve the weldability of the base material and clean the components before welding, the NGV was vacuum annealed like the IN738 alloy.

Given the small nominal thickness of the trailing edge (TE), a first dissimilar filler alloy, 3698-6B, was used to create a transition layer. A second dissimilar filler alloy, 3653BS, containing only 0.2 wt.% B and 1.2 wt.% Si, was then used to apply a high-strength and good oxidation-resistant layer. Due to the increased boron content in the top oxidation-resistant layer and the application of a protective aluminum coating to the NGV including the TE, a diffusion heat treatment was combined with post-weld annealing, primary and secondary aging into a single cycle, aiming to reduce costs and improve production efficiency.

The combined diffusion, annealing, and primary aging heat treatments are performed in a vacuum using the same standard heat treatment cycle as for IN738. A protective aluminum coating is applied via a secondary aging period of 24 hours at 845°C using a standard aluminum-based coating according to the relevant engine manual procedures. During this heat treatment, the aluminum diffuses into the base material, creating an aluminum coating that protects the base material and the repaired area from oxidation under operating conditions.

After repair, the NGV underwent standard FPI and dimensional inspection to ensure compliance with relevant repair standards. In addition to the above, samples for metallographic evaluation were extracted from the repaired area. Metallographic examination confirmed the FPI results, with no cracks found in the weld or HAZ. Figure 4 shows a micrograph showing the fusion line, base material, and weld.

Cracking in the HAZ and weld is eliminated by lowering the solidus temperature of the interdendritic eutectic formed in the weld during weld pool solidification to below the solidus temperature of IN738 (1201.45°C and 1284.53°C, respectively). This allows liquefaction cracks to self-heal along the fusion zone, as shown in Figure 1. Stress-strain cracks are eliminated by accommodating residual stresses in the transition layer, which has superior ductility compared to high-γ' IN738. Excellent oxidation resistance is achieved by optimizing the alloying elements in the top oxidation-resistant layer. The combined use of first and second dissimilar fillers allows welding at ambient temperatures, which distinguishes the preferred embodiment from the prior art, simplifies the technology, and reduces repair costs.

Claims (15)

1. Methods for repairing and manufacturing turbine engine components include the following steps: a) Preparing the base material for welding by removing damaged material and contaminants to expose a defect-free and uncontaminated base material; b) Applying a transition layer to a defect-free and uncontaminated matrix material via a fusion welding process and a first dissimilar filler, the first dissimilar filler comprising: (ix) at least one element selected from niobium, molybdenum, and tungsten, in a total amount of 2 wt.% to 25 wt.%. x) Nickel and balance impurities c) Diffusion heat treatment, which is performed for 30 minutes to 24 hours at a temperature exceeding the initial aging temperature but below the initial melting temperature of the base material after the application of the transition layer and before the application of the top oxidation-resistant layer. d) Apply a top oxidation-resistant layer to the transition layer using a fusion welding process and a second dissimilar filler, the second dissimilar filler comprising: e) Perform post-weld heat treatment selected from annealing, primary and secondary aging, and stress relief of the base material, or a combination thereof; and f) Restore the geometry of the repaired area. 2. The method for repairing and manufacturing turbine engine components according to claim 1, wherein the method includes an additional step of machining the transition layer. 3. The method for repairing and manufacturing turbine engine components according to claim 1 or 2, wherein the method includes an additional step of hot isostatic pressing prior to applying a transition layer. 4. The method for repairing and manufacturing turbine engine components according to claim 1 or 2, wherein the method includes an additional step of hot isostatic pressing after applying a transition layer. 5. The method for repairing and manufacturing turbine engine components according to claim 1 or 2, wherein the method includes an additional step of hot isostatic pressing using parameters selected from those defined for the base material after the application of the top oxidation-resistant layer but before the aging heat treatment. 6. The method for repairing and manufacturing turbine engine components according to claim 3, wherein the method includes an additional step of annealing heat treatment in a vacuum or protective atmosphere after hot isostatic pressing but before applying a transition layer. 7. The method for repairing and manufacturing turbine engine components according to claim 1 or 2, wherein the welding process is carried out at ambient temperature. 8. The method for repairing and manufacturing turbine engine components according to claim 1 or 2, wherein the welding process is performed by preheating the turbine engine components to a temperature of 600°C to 1100°C. 9. The method for repairing and manufacturing turbine engine components according to claim 1 or 2, wherein the matrix material of the engine component is selected from single crystal, directionally solidified, equiaxed nickel, cobalt and iron-based superalloys. 10. The method for repairing and manufacturing turbine engine components according to claim 1 or 2, wherein the turbine engine components are selected from nozzle guide vanes, compressor stationary vanes, compressor blades, high-pressure compressor blades, high-pressure turbine blades, intermediate-pressure turbine blades, low-pressure turbine blades, shrouds, sealing sections, housings, guide vanes, combustion chambers, flame tubes, fuel nozzles, and manifolds of aerospace and industrial turbine engines. 11. The method for repairing and manufacturing turbine engine components according to claim 1 or 2, wherein the welding process is selected from laser welding, gas-shielded tungsten inert gas welding, plasma welding, micro-beam plasma welding, and electron beam welding. 12. The method for repairing and manufacturing turbine engine components according to claim 1, wherein step f) is performed by a method selected from machining, grinding and polishing. 13. The method for repairing and manufacturing turbine engine components according to claim 1 or 2, wherein the method includes a non-destructive testing step and a dimensional inspection step. 14. The method for repairing and manufacturing turbine engine components according to claim 6, wherein the protective atmosphere is hydrogen. 15. A turbine engine component, which is repaired and manufactured by the method of repairing and manufacturing a turbine engine component according to any one of claims 1-14.