Micro-crack electric shock regulation and control method for laser remanufactured nickel-based superalloy aircraft engine blade

Abstract

The invention discloses a micro-crack electric shock regulation and control method for a laser remanufactured nickel-based superalloy aircraft engine blade, which comprises the following steps of: carrying out composite heat treatment on the laser remanufactured aircraft engine blade, adjusting the structure property and exposing the defects of the aircraft engine blade; scanning the laser remanufactured aircraft engine blade after heat treatment, performing three-dimensional reconstruction on scanning data, calibrating three-dimensional coordinates of the defect based on the reconstructed scanning data, and acquiring geometric characteristic parameters of the defect; electric shock process parameters are calculated according to the geometrical characteristic parameters of the defects, and an electrode loading path is planned on the laser remanufactured aircraft engine blade; performing electric shock treatment on the laser remanufactured aircraft engine blade point by point on the basis of the planned path and the electric shock process parameters to obtain a repaired aircraft engine blade; the laser additive remanufacturing nickel-based high-temperature alloy blade is subjected to high-energy short-time electric shock treatment, targeted precise repair of internal micro-nano defects is achieved, and the service life of the blade is prolonged.

Description

Technical Field

[0001] The present invention relates to the technical field of laser additive repair, and particularly to a method for regulating micro-cracks of a nickel-based superalloy aero-engine blade by electro-impulse in laser remanufacturing. Background Art

[0002] The turbine blade of an aero-engine is a key component affecting the thrust-to-weight ratio and reliability of the engine, and its performance directly determines the service life and safety margin of the aero-engine. The turbine inlet temperature of modern aero-engines has exceeded 1800 °C. Under such extreme working conditions, the blade needs to withstand multi-field coupling effects such as a centrifugal load of up to 300 MPa, a non-uniform temperature field of about 1600 °C, and a rotational speed of 20000 r / min. This service environment with long-term high temperature (800 - 1100 °C), high pressure (2 - 3 MPa gas pressure), and strong mechanical load (equivalent to 20000 times the self-weight of the blade) causes complex failure behaviors on the blade surface: fretting wear occurs in the tip region due to high-speed gas erosion, local ablation is caused by the spalling of the thermal barrier coating at the blade edge, and fatigue cracks are initiated at the dovetail joint under the action of alternating stress. Statistical data from Pratt & Whitney in the United States shows that within a typical 2000-hour overhaul period, the local failure rate of turbine blades is as high as 78%, and the overall scrapped quantity due to local damage accounts for 92% of the total replacement quantity, and the actual effective life only lasts for 500 - 1000 hours.

[0003] At present, aviation turbine blades are mainly manufactured using third-generation nickel-based single-crystal superalloys (such as CMSX-4, DD6, etc.), and their material cost is as high as $3000 - 5000 / kg. These alloys form γ' strengthening phases by adding elements such as 8 - 12% Cr, 5 - 7% Co, and 3 - 5% Al, and possess excellent high-temperature strength and oxidation resistance. However, the manufacturing of directionally solidified single-crystal blades requires a complex mold shell system and precise temperature gradient control, and the yield rate is less than 40%. The repairable area in failed blades accounts for more than 65%. Traditional repair processes such as gas tungsten arc welding (GTAW) have problems such as matrix recrystallization and heat-affected zone (HAZ) softening due to large heat input (>500 J / mm), while plasma spraying (PS) cannot meet the requirements of aerodynamic loads due to low bonding strength (<200 MPa). Laser Additive Remanufacturing (LAR) technology uses a coaxial powder feeding system to layer-by-layer cladding (layer thickness 50 - 100 μm) at the defective part, and uses a high-energy laser beam (power 1 - 3 kW, spot diameter 0.3 - 0.8 mm) to achieve rapid melting and solidification in a micro-region (<1 mm3) (cooling rate 103 - 105 K / s). Combining online topography detection (accuracy ±0.05 mm) and adaptive path planning, the repair cycle can be shortened to 48 hours, and the cost is only 18 - 22% of that of newly manufactured parts. However, the high γ' phase content (>60%) of nickel-based alloys results in a room-temperature elongation rate of only 4 - 6%. During the process of rapid laser melting and solidification (temperature gradient up to 106 K / m), due to elemental microsegregation (such as the enrichment of Ta and Re between dendrites), a eutectic reaction (L→γ+γ') is triggered, forming network-shaped grain boundary carbides (M23C6 type) in the repair area. Coupled with the tensile stress generated by non-equilibrium solidification shrinkage (peak value up to 800 MPa), the crack sensitivity index (CST) exceeds 4.5, far higher than the acceptable threshold of 2.0. CT detection by the Fraunhofer ILT in Germany shows that the internal defect density of the repair layer reaches 32 / mm3 (porosity >1.2%), and the crack growth rate (da / dN) is 3 - 5 times higher than that of the matrix, reducing the low-cycle fatigue life of the repaired blades to 57 - 63% of that of newly manufactured blades, which has become the key bottleneck restricting the application of laser remanufacturing technology in failed aero-engine turbine blades. At present, research on eliminating defects in nickel-based superalloy additive remanufacturing of aero-engine blades mainly focuses on crack suppression during the additive process. For example, through process parameter optimization, alloy composition fine-tuning, matrix preheating, etc., macro-cracks (≥20 μm) in the additive process can be basically suppressed, but micro-cracks (<20 μm) are difficult to eradicate due to the intrinsic characteristics of materials and processes. In addition, during the post-heat treatment process of remanufactured blades, due to the rapid precipitation of high-volume fraction γ' strengthening phases (50 - 70%) in nickel-based superalloys, stress release mismatch leads to the initiation of new cracks and the further propagation of original residual micro-cracks.Hot isostatic pressing can partially eliminate tiny internal defects in materials, but the entire material needs to be heated above the solid solution temperature of the material (1100 - 1200 °C). The high-temperature environment accelerates grain boundary migration and grain growth, leading to tissue coarsening, reducing the creep strength of the material, and having high energy consumption and a long cycle (10 - 30 hours).

[0004] Therefore, it is necessary to provide a method for regulating micro-crack electroshock in laser remanufactured nickel-based superalloy aero-engine blades to achieve "targeted" repair of defects and extend the service life of the blades. Summary of the Invention

[0005] In view of this, the present application provides a method for regulating micro-crack electroshock in laser remanufactured nickel-based superalloy aero-engine blades, which is used to solve the problem of extending the service life of laser remanufactured nickel-based superalloy aero-engine blades.

[0006] To achieve the above technical objectives, the present application adopts the following technical solutions:

[0007] In the first aspect, the present application provides a method for regulating micro-crack electroshock in laser remanufactured nickel-based superalloy aero-engine blades, including the following steps:

[0008] Perform composite heat treatment on the laser remanufactured aero-engine blade to adjust the tissue properties and expose defects;

[0009] Scan the laser remanufactured aero-engine blade after heat treatment, three-dimensionally reconstruct the scanned data, and calibrate the three-dimensional coordinates of the defects and quantify the geometric feature parameters of the defects based on the reconstructed data;

[0010] Calculate the electroshock process parameters according to the geometric feature parameters of the defects and establish an electrode loading path on the laser remanufactured aero-engine blade;

[0011] Load the electrodes point by point on the laser remanufactured aero-engine blade according to the path and electroshock process parameters and perform electroshock treatment to obtain the repaired aero-engine blade.

[0012] Preferably, the process of performing composite heat treatment on the laser remanufactured aero-engine blade is as follows: stress relief annealing of the laser remanufactured aero-engine blade at 550 - 650 °C, then solution heat treatment at 1150 - 1200 °C, and then aging heat treatment at 730 - 800 °C; and / or, the heating and cooling rates of stress relief annealing are 1 - 3 °C / min.

[0013] Preferably, the specific steps for scanning the heat-treated laser remanufactured aero-engine blade are as follows: performing three-dimensional tomography on the heat-treated laser remanufactured aero-engine blade and performing dynamic exposure compensation; performing three-dimensional tomography on the core area of the laser remanufactured aero-engine blade at a resolution of 0.5 μm, and performing three-dimensional tomography on the non-core area of the laser remanufactured aero-engine blade at a resolution of 2 μm; the core area of the laser remanufactured aero-engine blade is the laser additive remanufacturing forming area, and the non-core area of the laser remanufactured aero-engine blade is the interface bonding area and the heat-affected area where the blade substrate is connected to the laser additive remanufacturing forming area.

[0014] Preferably, the steps for quantifying the geometric feature parameters of the defect are as follows: generating a three-dimensional point cloud model of the defect through voxelization processing and surface reconstruction algorithms, and quantifying the geometric feature parameters of the microcrack; the geometric feature parameters include crack length L, width W, depth D, and aspect ratio L / W.

[0015] Preferably, the electroshock process parameters include shock current density, pulsed current frequency, shock action time, and number of shocks;

[0016] The calculation formula for the shock current density is shown in Equation (Ⅰ):

[0017]

[0018] where J is the current density (A / mm 2 ), V d is the defect volume (mm 3 ), L d , W d are the defect length and width (mm), n is the defect shape factor, σ y is the material yield strength (MPa), ν is the Poisson's ratio, and k1 is the material conductivity correction coefficient;

[0019] The calculation formula for the pulsed current frequency is shown in Equation (Ⅱ):

[0020]

[0021] where f is the pulse frequency (Hz), d is the defect depth (mm), δ is the thickness of the thermal diffusion layer (mm), α is the material thermal conductivity (W / mK), D is the conductivity (S / m), c is the material specific heat capacity (J / kg / K), and k2, k3 are the process correction coefficients;

[0022] The calculation formula for the shock action time is shown in Equation (Ⅲ):

[0023]

[0024] where t is the single-point action time (s), ε eis the critical strain threshold, E is the elastic modulus (GPa), η is the energy conversion efficiency, and Ad / A0 is the proportion of the defect cross-section;

[0025] The calculation formula for the number of impacts is shown in Equation (Ⅳ):

[0026]

[0027] where N is the number of impacts, Creg is the target crack closure, C0 is the initial closure, △C is the closure increment per impact, and k5 is the material response coefficient.

[0028] Preferably, the specific steps for establishing the electrode loading path on the laser remanufactured aero-engine blade are as follows: longitudinally arrange the positive and negative electrodes along the leading edge - trailing edge direction of the blade for longitudinal movement, and then transversely arrange the positive and negative electrodes along the convex edge - concave edge direction of the blade for transverse movement. The distance between adjacent impact points is 3 - 5 mm.

[0029] Preferably, the process of electro-impulse treatment is as follows: real-time monitor the temperature T of the impact area through an infrared thermal imager. When T ≥ 0.8Tm, terminate the impact at the current point. After natural cooling to T ≤ 50 °C, perform the operation at the next impact point, where Tm is the γ' phase transformation temperature of the substrate.

[0030] Preferably, it further includes the steps of: finishing the repaired aero-engine blade, detecting the residual defect rate and resistivity, and calculating the blade life. When the blade life reaches 90% of the life of the new product of the same model, it is determined that the regulation is qualified; the finishing includes milling and surface polishing; the surface roughness Ra value of the repaired aero-engine blade after finishing is ≤ 0.4 μm, and the Ra value in the key aerodynamic profile area is ≤ 0.2 μm.

[0031] Preferably, the residual defect rate is detected by CT scanning, and the resistivity is detected by the four-probe method.

[0032] Preferably, the calculation formula for the blade life is shown in Equation (Ⅴ):

[0033] L = η · (σ e / σ0) 3.2 × [1 + erf((R c -R) / (0.15R c ))] × exp(-2.5D 0.6 ) Equation (Ⅴ),

[0034] where L is the predicted life (h), η is the process reliability coefficient, σe is the equivalent yield strength (MPa), σ0 is the yield strength of the original material, R is the measured resistivity (μΩ·m), Rc is the critical resistance threshold, T is the environmental temperature condition for the predicted life, D is the volume defect rate (%), and Dc is the maximum allowable defect rate.

[0035] The beneficial effects of this application are as follows: By performing high-energy short-time electric shock treatment on the laser additive remanufactured nickel-based superalloy blade, the targeted and precise repair of internal micro-nano defects is achieved, and the service life of the blade is extended.

[0036] Using the electric shock regulation method for microcracks of laser remanufactured nickel-based superalloy aeroengine blades in this application, compared with the traditional isostatic pressing method, the electric shock can be precisely focused on the defect area, avoiding the coarsening effect of overall heating on other normal areas. In addition, the short action time from milliseconds to minutes significantly improves the processing efficiency, enhances the strength, toughness and fatigue performance of the remanufactured aeroengine turbine blades, and further extends the service life of the remanufactured nickel-based superalloy aeroengine blades. BRIEF DESCRIPTION OF THE DRAWINGS

[0037] Figure 1 It is the electric shock targeted repair mechanism of this application;

[0038] Figure 2 It is the electrode loading path;

[0039] Figure 3 It is the process flow chart of this application. DETAILED DESCRIPTION OF THE EMBODIMENTS

[0040] In order to make the objectives, technical solutions and advantages of the present invention clearer, the present invention will be further described in detail below with reference to the embodiments. It should be understood that the specific embodiments described herein are only used to explain the present invention and are not used to limit the present invention.

[0041] This application provides an electric shock regulation method for microcracks of laser remanufactured nickel-based superalloy aeroengine blades, including the following steps:

[0042] S1. Composite heat treatment of the laser remanufactured aeroengine blade to adjust the tissue performance and expose the defects;

[0043] S2. Scan the heat-treated laser remanufactured aeroengine blade, three-dimensionally reconstruct the scan data, and calibrate the three-dimensional coordinates of the defects and quantify the geometric feature parameters of the defects based on the reconstructed data;

[0044] S3. Calculate the electric shock process parameters according to the geometric feature parameters of the defects and establish an electrode loading path on the laser remanufactured aeroengine blade;

[0045] S4. Load the electrodes point by point on the laser remanufactured aeroengine blade according to the path and the electric shock process parameters and perform electric shock treatment to obtain the repaired aeroengine blade.

[0046] This application realizes the targeted and precise repair of internal micro-nano defects by performing high-energy short-time electric shock treatment on the laser additive remanufactured nickel-based superalloy blade. As Figure 1As shown in the figure, its working mechanism is as follows: When a pulsed current passes through a metal material containing defects (such as cracks and pores), due to the "flow-around effect", the current density near the defects is high, which will generate instantaneous local high temperature, passivate or even melt the crack tip, and inhibit crack propagation. At the same time, the non-uniform temperature field generates thermal compressive stress, promoting the contact closure of the crack surface; in addition, the electromigration effect of the electromagnetic pulse can also enhance atomic diffusion and dislocation slip, further filling the crack, thereby achieving the "targeted" repair of the defect.

[0047] In some embodiments, the process of performing composite heat treatment on the laser remanufactured aeroengine blade is as follows: The laser remanufactured aeroengine blade is stress-relieved annealed at 550 - 650 °C, then solution heat-treated at 1150 - 1200 °C, and then aged heat-treated at 730 - 800 °C; and / or, the heating and cooling rates of the stress-relieving annealing are 1 - 3 °C / min.

[0048] In the present application, the specific process of step S1 is as follows: Stress-relieving annealing, heating the laser remanufactured blade to 550 - 650 °C at a rate of 1 - 3 °C / min, holding for 2 - 4 hours and then slowly cooling at the same rate to eliminate residual stress and avoid stress concentration; Solution heat treatment, uniformly heating the annealed blade to 1050 - 1150 °C, holding for 2 - 4 hours to fully dissolve aluminum, cobalt, and chromium elements in the γ matrix to form a solid solution with uniform composition; Ageing heat treatment, heating the solution-treated blade to 730 - 800 °C, holding for 10 - 20 hours (dynamically adjusted according to the blade size and alloy type) to promote the precipitation of γ' strengthening phase; The composite heat treatment achieves a triple synergistic effect: eliminating residual stress and exposing microcracks through gradient temperature control; enhancing the high-temperature strength and oxidation resistance of the matrix; optimizing the creep resistance.

[0049] In some embodiments, the specific steps of scanning the heat-treated laser remanufactured aeroengine blade are as follows: Performing three-dimensional tomography on the heat-treated laser remanufactured aeroengine blade and performing dynamic exposure compensation; The core area of the laser remanufactured aeroengine blade is scanned in three dimensions with a resolution of 0.5 μm, and the non-core area of the laser remanufactured aeroengine blade is scanned in three dimensions with a resolution of 2 μm; The core area of the laser remanufactured aeroengine blade is the laser additive manufacturing forming area, and the non-core area of the laser remanufactured aeroengine blade is the interface bonding area and the heat-affected area where the blade matrix is connected to the laser additive manufacturing forming area.

[0050] In some embodiments, the steps of quantifying the geometric feature parameters of the defect are as follows: Generating a three-dimensional point cloud model of the defect through voxelization processing and surface reconstruction algorithms to quantify the geometric feature parameters of the microcrack; The geometric feature parameters include crack length L, width W, depth D, and aspect ratio L / W.

[0051] Step S2 of this application is CT scan defect location and reconstruction in the remanufacturing area, that is, digital characterization of defects; specifically including: CT scan and data acquisition, using a high-resolution industrial X-ray CT scanner to perform multi-angle layer-by-layer scans on the remanufacturing area of the aero-engine turbine blade, and combining dynamic exposure compensation technology to eliminate geometric artifacts; multi-scale scanning strategy: scanning the core area of the blade (≤1 mm from the surface) at a resolution of 0.5 μm and the edge area (>1 mm) at a resolution of 2 μm to balance scanning efficiency and accuracy; three-dimensional reconstruction and defect location: through multi-scan data fusion and three-dimensional reconstruction technology, converting two-dimensional CT data into three-dimensional point clouds, and calibrating the three-dimensional coordinates and size parameters (length L, width W, depth D) of defects such as cracks and pores; defect morphology quantification: based on voxelization processing and surface reconstruction algorithms, using elliptical / rectangular geometric models to fit the defect morphology, and calculating key parameters such as aspect ratio (L / W), volume ratio (Vdefect / Varea); automatic defect recognition: introducing automatic detection algorithms based on gray threshold segmentation, region growing, and machine learning to improve defect detection efficiency and objectivity. Among them, gray threshold segmentation extracts the area with a gray value lower than 15-25% of the substrate average value as candidate defects; the region growing algorithm expands the candidate defects to the area where the gray difference of adjacent voxels ≤5%; the convolutional neural network classifier discriminates the authenticity of the candidate defects and outputs the probability values of cracks, pores, or artifacts. The digital characterization of defects provides accurate defect data support for subsequent electro-shock process planning, where the defect location accuracy in the core area ≤1 μm and in the non-core area ≤5 μm.

[0052] In some embodiments, the electro-shock process parameters include shock current density, pulsed current frequency, shock action time, and number of shocks;

[0053] The calculation formula for the shock current density is shown in Equation (Ⅰ):

[0054]

[0055] Where, J is the current density (A / mm 2 ), V d is the defect volume (mm 3 ), L d , W d are the defect length and width (mm), n is the defect shape factor. When L d / W d ∈[0.8, 1.2] (approximate circular defect), the shape factor is 1. When (strip crack), the shape factor is increased to 1.9, σ y is the material yield strength (MPa), ν is the Poisson's ratio, and k1 is the material conductivity correction coefficient (0.8 - 1.2);

[0056] The calculation formula for the pulse current frequency is shown in Equation (Ⅱ):

[0057]

[0058] Where f is the pulse frequency (Hz), d is the defect depth (mm), δ is the thickness of the thermal diffusion layer (mm), α is the thermal conductivity of the material (W / mK), D is the conductivity (S / m), c is the specific heat capacity of the material (J / kg / K), k2 and k3 are process correction factors; the exponential term controls the thermal penetration depth, and for deep defects, the frequency needs to be reduced to prevent surface overheating; the frequency is inversely proportional to the square of the thermal diffusion to ensure that the number of thermal cycles in the defect area meets the requirements of dislocation migration;

[0059] The calculation formula for the impact action time is shown in Equation (Ⅲ):

[0060]

[0061] Where t is the single-point action time (s), ε e is the critical strain threshold, E is the elastic modulus (GPa), η is the energy conversion efficiency, Ad / A0 is the proportion of the defect cross-section; the time is positively correlated with the square root of the defect cross-sectional area, and larger-sized defects require longer time to complete plastic deformation and diffusion fusion; the numerator represents the material's resistance to deformation, and the denominator reflects the electrical energy input efficiency;

[0062] The calculation formula for the number of impacts is shown in Equation (Ⅳ):

[0063]

[0064] Where N is the number of impacts, Creg is the target crack closure (95%), C0 is the initial closure (CT detection value), △C is the single-impact closure increment (0.2 - 0.4), k5 is the material response coefficient, ceil() represents rounding down, and an exponential decay model is used to predict the cumulative effect.

[0065] In some embodiments, the specific steps for establishing the electrode loading path on the laser remanufactured aero-engine blade are as follows: longitudinally arrange positive and negative electrodes along the leading edge - trailing edge direction of the blade for longitudinal movement, and then transversely arrange positive and negative electrodes along the convex edge - concave edge direction of the blade for transverse movement, and the spacing between adjacent impact points is 3 - 5 mm.

[0066] In this application, step S3 is the dynamic electric shock planning for the remanufacturing area, which specifically includes the multi-point electric shock position planning and the process parameter planning. First, according to the remanufacturing position dimensions at the blade tip, the positive and negative electrodes are respectively loaded at the leading and trailing edges of the blade for longitudinal shock. Then, in the transverse direction of the blade, the positive and negative electrodes are respectively loaded at the convex and concave edges of the blade, keeping the distance between the positive and negative electrodes the closest. After one point of shock, the positive and negative electrodes move synchronously to the next point, and the distance between adjacent two shock points is less than 5 mm, and the shock points in the transverse direction are evenly distributed to shock the entire transverse side length. According to the detected defect size, the shock current density, frequency, action time, and shock times are quantitatively calculated to achieve precise control of the shock parameters for different positions and different defect sizes.

[0067] In some embodiments, the process of the electric shock treatment is as follows: The temperature T of the shock area is monitored in real time by an infrared thermal imager. When T≥0.8Tm, the shock at the current point is terminated, and the operation of the next shock point is carried out after natural cooling to T≤50°C, where Tm is the γ' phase transformation temperature of the base material.

[0068] In this application, step S4 is the closed-loop electric shock treatment and monitoring, which specifically includes shock parameter setting, point-by-point shock treatment, online monitoring of the temperature change during the shock process, and process control. According to the planned process route and the calculated shock parameters for each point, the shock electrode is loaded at the main position, and then the corresponding parameters are set for shock treatment. During the shock process, the temperature change of the shock point is monitored in real time by an infrared thermal imager. When the temperature exceeds 0.8Tm, the shock at the current position ends, where Tm is the γ' phase transformation temperature of the superalloy material, and the shock of the next point does not start until the temperature drops to 50°C.

[0069] In some embodiments, it further includes the step: S5. Finish machining the repaired aero-engine blade, detect the residual defect rate and resistivity, and calculate the blade life. When the blade life reaches 90% of the life of the new product of the same model, it is determined that the regulation is qualified; the machining includes milling and surface polishing; the surface roughness Ra value of the repaired aero-engine blade after finish machining is ≤0.4 μm, and the Ra value in the key aerodynamic profile area is ≤0.2 μm.

[0070] In some embodiments, the residual defect rate is detected by CT scanning, and the resistivity is detected by the four-probe method.

[0071] In some embodiments, the calculation formula of the blade life is shown in formula (Ⅴ):

[0072] L=η·(σ e / σ0) 3.2 ×[1+erf((R c -R) / (0.15R c ))]×exp(-2.5D 0.6 ) Formula (Ⅴ),

[0073] Wherein, L is the predicted life (h), η is the process reliability coefficient (0.85 - 1.15), σe is the equivalent yield strength (MPa), σ e = σ y ·[1 - 0.8(D / D c ) 1.5 , σ0 is the yield strength of the original material, R is the measured resistivity (μΩ·m), Rc is the critical resistance threshold, R C = 1.25R0 + 0.03(T - 800), T is the environmental temperature condition of the predicted life, D is the volume defect rate (%), Dc is the maximum allowable defect rate (1.8%), erf is the error function, characterizing the contribution of the resistance parameter to the defect closure.

[0074] In this application, step S5 is the machining and quality evaluation of the remanufactured aero-engine blade. Specifically, it includes milling, surface polishing, statistically analyzing the residual defect rate through CT scanning, combining resistance measurement to predict the remanufacturing life of the blade, performing finish machining on the blade after electroshock treatment according to the precise CAD model of the blade by a five-axis linkage machine tool, then performing surface polishing on the grinding head, scanning the defects in the final remanufacturing area of the blade again through X-ray CT, calculating the defect rate, measuring the resistivity of the remanufacturing area of the blade, and predicting the life after repair through the formula to achieve remanufacturing quality evaluation. In addition, surface fine polishing is carried out by means of grinding head polishing, and the surface roughness of the blade is gradually reduced through multi-stage abrasives and fine polishing paste; the four-probe method is used to measure the resistivity of the remanufacturing area of the blade. The four-probe method can effectively eliminate the influence of contact resistance and provide accurate material resistivity values. Multiple representative positions are selected in the remanufacturing area of the blade for measurement, and the average value is taken as the resistivity of this area.

[0075] The following further illustrates this solution through specific embodiments.

[0076] Source of raw materials for laser remanufactured aero-engine blade: For the first-stage turbine blade of a certain type of aero-engine, the blade material is the third-generation nickel-based single-crystal superalloy (CMSX-4). After the service life reaches 1500 hours, slight ablation and microcracks appear in the tip region (the average depth is about 20μm, and the maximum depth is less than 30μm)

[0077] Example 1

[0078] A method for electroshock regulation of microcracks in a laser remanufactured nickel-based superalloy aero-engine blade, comprising the following steps:

[0079] S1. Place the laser remanufactured aeroengine blade in a vacuum annealing furnace, heat it to 600 °C at a heating rate of 2 °C / min, hold for 3 hours, and then cool it to room temperature at a cooling rate of 1.5 °C / min to complete stress relief annealing; then continue to place the annealed blade in a vacuum furnace, heat it to 1100 °C, hold for 2.5 hours, and then air cool at 30 °C / min to complete solution heat treatment; place the solution heat treated blade in an aging furnace, heat it to 760 °C, hold for 15 hours, and then air cool to room temperature to complete aging heat treatment;

[0080] S2. Use the YXLON FF20 industrial CT scanner of Fraunhofer ILT in Germany to scan the remanufactured area at the blade tip after heat treatment. The scanning parameters are: X-ray voltage 180 kV, current 150 μA, and spatial resolution 5 μm. Adopt the multi-angle helical scanning method, and the scanning angle range is 360°; then, use VGStudio MAX software to perform 3D reconstruction on the CT scan data, adopt the algorithm based on gray threshold segmentation to identify crack and pore defects, and measure their sizes and positions;

[0081] S3. Fix the blade on the electroshock platform, connect the positive electrode to the leading edge of the blade (near the crack position), and the negative electrode to the trailing edge of the blade. Then connect the positive electrode to the convex edge of the blade (near the crack position), and the negative electrode to the concave edge of the blade. As Figure 2 shown, plan a series of impact points in the laser remanufactured repair area, with the spacing between adjacent impact points being 3 mm, evenly distributed along the crack length direction;

[0082] According to formula (Ⅰ), the yield strength of CMSX-4 alloy is 800 MPa, the Poisson's ratio is 0.33, the material conductivity correction coefficient is 1.1, the defect volume is 6000 μm 3 , the defect length is 60 μm, the width is 5 μm, the aspect ratio is 12, and the morphology correction coefficient n is taken as 1.9 (strip crack), and the calculated current density is 65 A / mm 2 ;

[0083] According to formula (Ⅱ), the thermal conductivity of CMSX-4 alloy is 29 W / mK, the electrical conductivity is 1.2×10^6 S / m, the specific heat capacity is 420 J / kg / K, the defect depth is 20 μm, the process correction coefficients k2 = 0.5, k3 = 1.2, and the calculated pulse frequency is 78 Hz. Among them, δ is the thickness of the thermal diffusion layer, taken as the defect depth d = 20 μm = 20×10 -6 m, the density of CMSX-4 ρ≈8700 kg / m 3 ;

[0084] According to formula (Ⅲ), with a critical strain of 0.2, an elastic modulus of 200 GPa, an energy conversion efficiency of 0.7, a proportion of the defect cross-section of 0.1, and a material resistivity of 0.83×10^-6 Ω·m, the single impact time is calculated to be 0.02 m;

[0085] According to formula (Ⅳ), with a target crack closure of 95%, an initial closure of 20% (estimated based on CT scan results), a single impact closure increment of 30%, and a material response coefficient of 0.01, the number of electric shocks is calculated to be 5 times;

[0086] S4. Use the Pulsotron E500 type electric pulse device of Pulsotronic, Germany, to control the current, voltage, frequency, and pulse width. Set the current density to approximately 65 A / mm 2 , the pulse frequency to 78 Hz, the single impact time to 20 ms, and the number of electric shocks to 5 times. Align the electrode with the first impact point and start the electric pulse device for impact. Perform impact treatment point by point along the crack length direction to ensure that the impact area covers the entire crack area. Use a FLIR A655sc infrared thermal imager to monitor the temperature change of the impact point in real time. When the temperature exceeds 910 °C (the γ' phase transformation temperature of CMSX-4 alloy is approximately 1350 °C, and 0.7 is taken as the safety factor), immediately stop the impact at the current position. After the temperature drops below 50 °C, then proceed with the treatment of the next impact point;

[0087] S5. Use a five-axis CNC milling machine (DMG MORIDMU 50) to finish machine the blade, remove the possible surface oxide layer after electric shock treatment, and restore the aerodynamic shape of the blade. Control the machining accuracy within the dimensional tolerance of ±0.02 mm, and control the geometric tolerance of key parts within 0.01 mm; Use grinding heads with grit sizes of 400#, 800#, and 1200# to polish the blade surface in sequence, and then use alumina polishing paste for polishing, finally making the surface roughness Ra value of the blade ≤ 0.2 μm;

[0088] The industrial CT scanner was used again to scan the blade tip region to evaluate the defect rate after electroshock treatment. The results showed that the crack length decreased to nearly disappear, the width completely closed, and the volume defect rate decreased to less than 0.1%; the resistivity of the blade repair area was measured by the four-probe method, and the measured value was 0.84 μΩ·m, which was close to the resistivity of the matrix material, indicating good metallurgical bonding in the repair area; according to formula (Ⅴ), the process reliability coefficient was 0.98, the equivalent yield strength was 860 MPa, the yield strength of the original material was 820 MPa, the critical resistance threshold was 0.9 μΩ·m, the ambient temperature was 900 °C, and the allowable maximum defect rate was 1.8%. Then, it was predicted that the service life of the repaired blade was increased to more than 2,200 hours, indicating that the electroshock regulation method effectively improved the service life of the remanufactured blade.

[0089] This embodiment shows that the electroshock regulation method for microcracks of nickel-based superalloy aeroengine blades by laser remanufacturing can effectively repair the microcrack defects of the blades and significantly improve the performance and life of the remanufactured blades. The calculated current density and frequency are both within a reasonable range, verifying the feasibility and effectiveness of the method of the present invention.

[0090] The above is only a preferred specific embodiment of the present invention, but the protection scope of the present invention is not limited thereto. Any changes or substitutions that can be easily thought of by those skilled in the art within the technical scope disclosed by the present invention should be covered by the protection scope of the present invention.

Claims

1. A method for controlling microcracks of a laser remanufactured nickel-based high-temperature alloy aeroengine blade by electric shock, characterized in that: The following steps are involved: Composite heat treatment of laser remanufactured aeroengine blades to adjust their organizational properties and expose their defects; Scanning the laser remanufactured aeroengine blade after heat treatment, three-dimensionally reconstructing the scan data, calibrating the three-dimensional coordinates of the defect based on the reconstructed scan data and obtaining the geometric characteristic parameters of the defect; Calculating electric shock process parameters according to the geometric characteristic parameters of the defect, and planning an electrode loading path on the laser remanufactured aeroengine blade; Based on the planned path and the electric shock process parameters, electric shock treatment is performed point by point on the laser remanufactured aero-engine blade, so that the aero-engine blade is repaired.

2. The method for controlling microcracks of nickel-based high-temperature alloy aeroengine blades by electric shock during laser remanufacturing according to claim 1 is characterized in that: The process of composite heat treatment of the laser remanufactured aero-engine blade is as follows: stress relief annealing of the laser remanufactured aero-engine blade at 550-650°C, solution heat treatment at 1150-1200°C, and aging heat treatment at 730-800°C; and / or, the heating and cooling rates of the stress relief annealing are 1-3°C / min.

3. The method for controlling microcracks of nickel-based high-temperature alloy aeroengine blades by electric shock during laser remanufacturing according to claim 1 is characterized in that: The specific steps of scanning the laser remanufactured aero-engine blade after heat treatment are as follows: performing a three-dimensional tomography scan on the laser remanufactured aero-engine blade after heat treatment, and performing dynamic exposure compensation; performing a three-dimensional tomography scan on the core area of ​​the laser remanufactured aero-engine blade at a resolution of 0.5μm, and performing a three-dimensional tomography scan on the non-core area of ​​the laser remanufactured aero-engine blade at a resolution of 2μm; the core area of ​​the laser remanufactured aero-engine blade is the laser additive remanufacturing forming area, and the non-core area of ​​the laser remanufactured aero-engine blade is the interface bonding area and heat-affected zone where the blade substrate and the laser additive remanufacturing forming are disconnected.

4. The method for controlling microcracks of nickel-based high-temperature alloy aeroengine blades by electric shock during laser remanufacturing according to claim 1 is characterized in that: The step of quantifying the geometric characteristic parameters of the defect is: generating a three-dimensional point cloud model of the defect through voxelization processing and surface reconstruction algorithm, and quantifying the geometric characteristic parameters of the microcracks; the geometric characteristic parameters include crack length L, width W, depth D and aspect ratio L / W.

5. The method for controlling microcracks of nickel-based high-temperature alloy aeroengine blades by electric shock during laser remanufacturing according to claim 1 is characterized in that: The electric shock process parameters include shock current density, pulse current frequency, shock action time, and shock times; The calculation formula of the impact current density is shown in formula (I): Where J is the current density (A / mm 2 ), V d is the defect volume (mm 3 ), L d ,W d is the defect length and width (mm), n is the defect morphology factor, σ y is the material yield strength (MPa), ν is Poisson’s ratio, k1 is the material conductivity correction factor; The calculation formula of the pulse current frequency is shown in formula (II): Where, f is the pulse frequency (Hz), d is the defect depth (mm), δ is the thickness of the heat diffusion layer (mm), α is the thermal conductivity of the material (W / mK), D is the electrical conductivity (S / m), c is the specific heat capacity of the material (J / kg / K), k2 and k3 are process correction factors; The calculation formula of the impact action time is shown in formula (III): Where, t is the single point action time (s), ε e is the critical strain threshold, E is the elastic modulus (GPa), η is the energy conversion efficiency, A d / A0 is the defective cross-section ratio; The calculation formula of the impact times is shown in formula (IV): Where N is the number of impacts, Creg is the target crack closure, C0 is the initial closure, △C is the single impact closure increment, and k5 is the material response coefficient.

6. The method for controlling microcracks of nickel-based high-temperature alloy aeroengine blades by electric shock during laser remanufacturing according to claim 1, characterized in that: The specific steps of establishing the electrode loading path on the laser remanufactured aeroengine blade are as follows: the positive and negative electrodes are longitudinally arranged along the leading edge-trailing edge direction of the blade for longitudinal movement, and then the positive and negative electrodes are transversely arranged along the convex edge-concave edge direction of the blade for transverse movement, and the distance between adjacent impact points is 3-5mm.

7. The method for controlling microcracks of nickel-based high-temperature alloy aeroengine blades by electric shock during laser remanufacturing according to claim 1, characterized in that: The process of electric shock treatment is: real-time monitoring of the impact area temperature T by infrared thermal imager, termination of the current point impact when T≥0.8Tm, and execution of the next impact point operation after natural cooling to T≤50°C, where Tm is the γ' phase transition temperature of the substrate.

8. The method for controlling microcracks of nickel-based high-temperature alloy aero-engine blades by electric shock during laser remanufacturing according to claim 1 is characterized in that: It also includes the steps of: finishing the repaired aero-engine blade, detecting the residual defect rate and resistivity, and calculating the blade life, and the blade life is qualified if it reaches 90% of the life of a new blade of the same model; the finishing includes milling and surface polishing; the surface roughness Ra value of the repaired aero-engine blade after finishing is ≤0.4μm, and the Ra value of the key aerodynamic surface area is ≤0.2μm.

9. The method for controlling microcracks of nickel-based high-temperature alloy aeroengine blades by electric shock during laser remanufacturing according to claim 8, characterized in that: The residual defect rate is detected by CT scanning, and the resistivity is detected by a four-probe method.

10. The method for controlling microcracks of nickel-based high-temperature alloy aero-engine blades by electric shock during laser remanufacturing according to claim 8, characterized in that: The calculation formula of the blade life is shown in formula (V): L = η·(σ e / σ0) 3.2 ×[1 + erf((R c - R) / (0.15R c ))] × exp(-2.5D 0.6 ) Equation (V), Wherein, L is the predicted life (h), η is the process reliability coefficient, σe is the equivalent yield strength (MPa), σ0 is the yield strength of the original material, R is the measured resistivity (μΩ·m), Rc is the critical resistance threshold, T is the ambient temperature condition for the predicted life, D is the volume defect rate (%), and Dc is the maximum allowable defect rate.

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