A laser cladding repair method for turbine blade tip damage of an aero-engine

Abstract

The application relates to a laser cladding repair method for turbine blade tip damage of an aero-engine, and relates to the field of turbine blade repair, and comprises the following steps: S10, material inspection heat treatment is performed on a turbine blade, and it is confirmed whether a heterogeneous material repair area exists in a tip area; S20, flaw detection is performed on the turbine blade, and a crack position of the tip area is marked; S30, a crack part and a heterogeneous material of the tip area are removed; S40, the thickness of an exhaust edge in a distance range from a removed end face of the tip is measured, if the thickness is less than an original design thickness by a set amplitude, the exhaust edge is reinforced; S50, laser cladding is performed on the tip in a protective gas environment, and the removed part of the tip is recovered; S60, the cladding area of the tip is polished and shaped, and the original appearance of the tip is repaired; and S70, solid solution and aging treatment are performed on the blade, and the repair effect of the turbine blade tip damage of the aero-engine can be improved.

Description

Technical Field

[0001] This application relates to the field of turbine blade repair, and in particular to a method for repairing tip damage of aero-engine turbine blades by laser cladding. Background Technology

[0002] High-pressure turbine blades for aero engines gradually age due to a variety of complex factors during long-term operation, including high temperature, high pressure, alternating stress, airflow erosion, corrosion, and oxidation. As the turbine blades age, their performance and structural integrity continuously deteriorate, leading to a significant reduction in engine efficiency, structural failure, and endangering flight safety. High-pressure turbine blades mainly consist of four parts: the blade tip, the blade body, the blade spar, and the blade root. During aero-engine operation, the blade tip experiences the highest temperature and the most severe operating conditions. Cracks, wear, and other defects are mostly concentrated in the blade tip region. If not repaired in time, the cracks will gradually extend to the blade body, causing the blade to become completely unusable.

[0003] When defects appear at the tips of high-pressure turbine blades, repair is mainly carried out using two techniques: laser cladding and welding. Welding includes argon arc welding, plasma arc welding, and brazing, among others. Currently, foreign aero-engine manufacturers and repair companies such as MTU and GE increasingly use laser cladding for high-pressure turbine blade tip repair. Laser cladding delivers cladding material powder to the defect location at the blade tip through methods such as powder feeding, wire feeding, and powder spreading. A laser then scans and melts the powder according to a pre-set path and power, shaping the blade. Subsequent minor reshaping is all that's needed to complete the tip shaping. Compared to welding for repairing blade tip damage, laser cladding offers advantages such as lower heat input, lower thermal stress, more precise and faster shaping, less subsequent reshaping work, and lower repair costs.

[0004] Currently, the following problems still exist in the laser cladding repair of high-pressure turbine blade tips: For some blades whose tip areas have undergone dissimilar material welding, due to the difference in the thermal expansion coefficients of different materials, coupled with the low ductility of nickel-based high-temperature alloy materials, hot cracks are easily generated during recladding; the exhaust edge of some aged high-pressure turbine blade tips has a weak structure and is easily torn by the shrinkage stress generated by the cladding layer during the cladding process. Summary of the Invention

[0005] To improve the repair effect of turbine blade tip damage in aero-engines, this application provides a laser cladding repair method for turbine blade tip damage in aero-engines.

[0006] The laser cladding repair method for tip damage of aero-engine turbine blades provided in this application adopts the following technical solution: A laser cladding repair method for tip damage of an aero-engine turbine blade includes the following steps: S10, performing material inspection and heat treatment on the turbine blade to confirm whether there is a foreign material repair area in the tip region; S20, performing flaw detection on the turbine blade to mark the crack location in the tip region; S30, removing the cracked portion and foreign material from the tip region; S40, measuring the thickness of the exhaust edge within a set distance from the removed end face of the tip, and if the thickness is less than the original design thickness set value, reinforcing the exhaust edge; S50, performing laser cladding on the tip in a protective gas environment to restore the removed portion of the tip; S60, grinding and reshaping the clad area of ​​the tip to restore the original morphology of the tip; S70, performing solution treatment and aging treatment on the blade.

[0007] By employing the above technical solutions and utilizing the material inspection and heat treatment method for turbine blades, it is possible to reveal areas of dissimilar materials within the blade tip region. This allows for the determination of whether the turbine blade tip region has undergone dissimilar material repair based on the presence of visually distinct repaired areas of dissimilar materials, and also facilitates the removal of repaired dissimilar materials. The method of removing dissimilar materials within the blade tip region can prevent thermal cracking caused by differences in the coefficients of thermal expansion between different material regions during laser cladding. Furthermore, by reinforcing the exhaust edge with a thickness less than the original design thickness, it is possible to strengthen the structurally weak areas of the exhaust edge that develop after aging, preventing these weak areas from being torn by the shrinkage stress generated by the cladding layer.

[0008] In one specific implementation scheme, in step S10, the material inspection heat treatment method is as follows: the turbine blade is heated to 420-680°C at a heating rate of 20-50°C / min, held at that temperature for 0.2-1.8 hours, and then cooled to room temperature in the furnace.

[0009] By adopting the above technical solution, the turbine blades are heated to 420-680℃ at a heating rate of 20-50℃ / min, held at that temperature for 0.2-1.8 hours, and then cooled to room temperature in the furnace. This method allows different materials in the blade tip area to exhibit distinctly different visual appearances, which is helpful for determining whether the blade tip area has been repaired with foreign materials and for completely removing foreign repair materials.

[0010] In one specific implementation scheme, in step S20, the method for detecting flaws in the turbine blade includes fluorescent flaw detection, X-ray flaw detection, eddy current flaw detection, or penetrant flaw detection, and the detection content includes the number of cracks and the distribution range of cracks in the blade tip region.

[0011] By adopting the above technical solutions and using fluorescent testing, X-ray testing, eddy current testing, or penetrant testing methods to inspect turbine blades, the flexibility of turbine blade inspection can be improved. Detecting the number and distribution of cracks in the blade tip region facilitates the complete removal of cracks in this area, ensuring the effectiveness of laser cladding repair of the aero-engine turbine blade tip region.

[0012] In one specific implementation, in step S30, a small grinding device is used to remove all cracked areas, and if there are areas to be repaired with foreign materials, the foreign materials in the repair areas are also removed.

[0013] By adopting the above technical solution, all cracks and any foreign materials that may exist in the blade tip area can be removed using small grinding equipment. This makes the removal operation more convenient, ensures that there are no aging cracks in the blade tip area, and prevents the generation of new cracks during the repair process, thus guaranteeing the laser cladding repair effect.

[0014] In one specific implementation scheme, after step S30, the method further includes step S35: perform flaw detection on the turbine blade again, and if there is a crack in the blade tip area, execute step S30 again.

[0015] By adopting the above technical solution and conducting further flaw detection on the turbine blade after the cracked area has been removed, it can be ensured that there are no more cracks in the tip of the turbine blade after the cracked area has been removed, thus guaranteeing the structural performance of the remaining tip, preventing the generation of cracks in the repair area, and ensuring the repair effect of the turbine blade tip.

[0016] In one specific implementation, in step S40, the thickness of the exhaust edge within 1 cm of the removed end face of the blade tip is measured. If the minimum value of the exhaust edge thickness is less than 80% of the original design thickness, the exhaust edge is reinforced.

[0017] By adopting the above technical solution, and by reinforcing the part of the vent edge with insufficient thickness within 1 cm of the removed end face, it is possible to effectively resist the shrinkage stress of the cladding layer and prevent the shrinkage stress of the cladding layer from pulling into the vent edge and generating new cracks.

[0018] In one specific implementation, the method for reinforcing the exhaust edge is to form a reinforcing rib by laser cladding on the outer edge of the exhaust edge. The reinforcing rib extends 2-5 mm from the point of minimum thickness of the exhaust edge to both inner sides, and has a thickness of 0.5-0.9 mm.

[0019] By adopting the above technical solution, and using reinforcing ribs that extend 2-5mm from the point of minimum thickness on the exhaust edge to both inner sides with a thickness of 0.5-0.9mm, it is possible to effectively strengthen the weak areas of the exhaust edge structure. Furthermore, by reshaping the damaged area of ​​the blade tip after repair, the original shape of the blade tip area after repair can be guaranteed.

[0020] In one specific implementation, in step S50, before laser cladding the blade tip, the blade tip area is preheated by heat treatment. The heating rate of the heat treatment preheating is 20-50℃ / min, and the heating temperature is 400-700℃. After laser cladding the blade tip, the temperature is maintained for 2-10 minutes, and then cooled to room temperature at a cooling rate of 10-30℃ / min.

[0021] By adopting the above technical solution and utilizing heat treatment preheating of the blade tip region, the microstructure of the material in the laser cladding repair area can be effectively improved, ensuring the structural performance of the repair area.

[0022] In one specific implementation scheme, in step S50, when performing laser cladding on the blade tip, coaxial vision is used to generate the cladding path of the blade tip, an infrared thermometer is used to measure the temperature of the molten pool formed by laser cladding in real time, and the temperature of the molten pool is controlled to 1800-2400℃ by controlling the laser power. The laser power adjustment range is 200-500W, the powder feeding rate is 2-8g / min, the cladding speed is 1-6mm / s, and the height of the cladding layer formed after the molten pool solidifies is 2.5-4.0mm.

[0023] By adopting the above technical solution and using coaxial vision to generate the blade tip cladding path, the repair and shaping of the cladding material in the blade tip region can be guaranteed, ensuring sufficient allowance for subsequent shaping. Using an infrared thermometer to measure the temperature of the molten pool formed by laser cladding in real time allows for control of the laser power during cladding, keeping the molten pool temperature within the range of 1800-2400℃. This ensures the fusion and shaping of the cladding material and the substrate material, resulting in a cladding layer morphology close to the original blade tip morphology and reducing the amount of processing required in subsequent shaping steps.

[0024] In one specific implementation scheme, the protective gas environment is argon gas with a purity of 99.9% and a pressure of 1-2 MPa, and the water and oxygen content in the protective gas environment is below 20 ppm.

[0025] By adopting the above technical solution and using an argon protective gas environment with a purity of 99.9% and a pressure of 1-2 MPa, it is possible to achieve oxidation-free processing in the cladding area, promote the removal of inclusions and gases inside the cladding layer, and improve the quality of the cladding layer.

[0026] In summary, this application includes at least one of the following beneficial technical effects: By performing material inspection and heat treatment on turbine blades, the blade material in the blade tip area can present a different color visually compared to the different repair materials that may have been used in the previous repair process. This makes it easier to determine whether there are different repair materials in the blade tip area and facilitates the removal of different repair materials mixed in with the blade material. By measuring the thickness of the exhaust edge within a set distance range from the blade tip removal end face, and reinforcing both sides of the minimum thickness point of the exhaust edge that is less than the original design thickness setting value, the tensile stress generated by the solidification of the cladding material by the exhaust edge can be increased, preventing the weak areas of the turbine blade exhaust edge structure from cracking. By using laser cladding on the blade tip in a protective gas environment to restore the removed portion of the blade tip, the oxidation inclusions in the cladding layer material can be effectively avoided, the internal inclusions and gas in the cladding layer can be expelled, the quality of the cladding material repair layer can be improved, and the structural strength of the turbine blade tip after repair can be guaranteed.

[0027] Other features and advantages of this application will be described in detail in the following detailed description section. Attached Figure Description

[0028] Figure 1 This is a flowchart of one embodiment of the present application.

[0029] Figure 2 This is a visual image of the different materials in the tip region of a turbine blade after material inspection and heat treatment, according to one embodiment of this application. Figure 3 This is a schematic diagram showing the location of the reinforcing ribs in one embodiment of this application.

[0030] Figure 4 for Figure 3 Enlarged view of part A in the middle.

[0031] Explanation of reference numerals in the attached diagram: 1. Blade tip; 11. Dissimilar material area; 12. Exhaust edge; 13. Minimum thickness point; 14. Reinforcing rib; 15. Repair material structure; 2. Blade body; 3. Edge plate; 4. Blade root. Detailed Implementation

[0032] The specific embodiments of this application will now be described in detail with reference to the accompanying drawings. It should be understood that the specific embodiments described herein are for illustration and explanation only and are not intended to limit this application.

[0033] In the description of this application, it should be noted that, unless otherwise expressly specified and limited, the terms "set" and "connection" should be interpreted broadly. For example, they can refer to a fixed connection, a detachable connection, or an integral connection; they can refer to a direct connection or an indirect connection through an intermediate medium; they can refer to the internal communication of two elements or the interaction between two elements. Those skilled in the art can understand the specific meaning of the above terms in this application according to the specific circumstances.

[0034] An embodiment of the laser cladding repair method for tip damage of aero-engine turbine blades in this application is as follows: Figure 1 As shown, it includes the following steps: S10. Perform material inspection and heat treatment on the turbine blades to confirm whether there are any areas requiring repair with different materials in the blade tip region.

[0035] Because aero-engine turbine blades operate in harsh environments with high temperatures and pressures for extended periods, they are prone to aging and cracking during operation, typically undergoing multiple crack repairs throughout their service life. During crack repair, dissimilar repair materials may be introduced. If such materials are present in the blade, thermal cracks can easily occur during laser cladding repair due to differences in the coefficients of thermal expansion between the different materials, affecting the repair outcome. Therefore, it is necessary to determine the presence of dissimilar repair materials in the turbine blade before proceeding with repair, in order to address any such materials appropriately.

[0036] By performing material inspection and heat treatment on turbine blades, and controlling the process parameters of the heat treatment, different material areas can produce different visual appearances. This allows for the identification of dissimilar repair materials and their locations by observing areas with varying visual appearances on the turbine blades.

[0037] Alternatively, communication can be established with the OEM before turbine blade repair, allowing the OEM to determine, based on maintenance records, whether the turbine blade to be repaired is undergoing its first repair. For turbine blades that are not undergoing their first repair, it can be further determined whether dissimilar materials were used for repair in previous repairs. For turbine blades that have been repaired with dissimilar materials, this step is performed to determine the location and extent of the dissimilar repair material. For turbine blades undergoing their first repair or those that have not been repaired with dissimilar repair materials before, this step is skipped, and the operation in step S20 is performed directly.

[0038] S20. Perform flaw detection on the turbine blades and mark the location of cracks in the blade tip area.

[0039] Flaw detection of turbine blades can be performed using various existing methods. By inspecting turbine blades, the extent and location of cracks inside the blade tip can be identified. The location and extent of all cracks can be marked using a marker or suitable dye, facilitating subsequent removal of the cracked portion at the blade tip.

[0040] S30, Remove cracked areas and foreign materials from the blade tip region.

[0041] Using a cutting or grinding device, the area with crack markings at the blade tip and the area of ​​dissimilar material shown after material inspection and heat treatment are cut off or ground away. The cutting or grinding range usually extends more than 1 mm beyond the crack marking area or the area of ​​dissimilar material.

[0042] S40. Measure the thickness of the exhaust edge within the set distance range from the blade tip removal end face. If the thickness is less than the original design thickness setting value, then reinforce the exhaust edge.

[0043] The exhaust edge is located on the rear side of the turbine blade in the direction of rotation. During the operation of an aero-engine, the temperature at the exhaust edge of the turbine blade is usually higher, and it is more prone to aging and thinning during operation. If the exhaust edge ages to a certain extent, when laser cladding repair is performed on the blade tip, the thinned area close to the cladding processing position is easily torn by the shrinkage stress generated by the solidification of the cladding layer.

[0044] To prevent cracking in the weakened areas of the exhaust edge due to aging, this application measures the thickness of the exhaust edge within a set distance from the blade tip removal end to identify weak areas of the exhaust edge near the blade tip removal end. If the thickness of the exhaust edge area is found to be thinner than its original design thickness by more than a set value, posing a risk of cracking during the cladding process, then that area is reinforced.

[0045] Various suitable methods can be used to reinforce the vent edge, such as setting reinforcing ribs, reinforcing plates, or applying cured reinforcing adhesive to the weak area. The reinforcing material is used to counteract the tensile stress generated by the cladding layer and prevent cracks from forming in the weak area under tensile stress.

[0046] S50. Laser cladding is performed on the blade tip in a protective gas environment to restore the removed portion of the blade tip.

[0047] A protective gas is introduced at the tip removal end of the turbine blade, and laser cladding additive manufacturing is performed on the tip removal end face using laser cladding equipment. The removed tip portion is then recovered using the additive material. The additive material typically uses high-temperature alloy powder that is the same as the turbine blade substrate material. Specifically, the additive material is customized according to the composition of the turbine blade substrate to obtain a nickel-based high-temperature alloy with the same composition as the turbine blade substrate. This nickel-based high-temperature alloy is then powdered using a rotating electrode to obtain nickel-based high-temperature alloy powder with a particle size of 20-200 μm.

[0048] Nickel-based superalloy powder is fed onto the blade tip removal surface and melted under laser irradiation to form a molten pool. After cooling, the molten pool solidifies on the blade tip removal surface, repairing the removed blade tip material. By controlling the parameters of laser cladding additive manufacturing, nickel-based superalloy material is deposited on the blade tip removal surface, forming a cladding layer structure that can cover the original morphology of the blade tip.

[0049] S60. Grind and reshape the cladding area of ​​the leaf tip to restore the original shape of the leaf tip.

[0050] Based on the original morphology data of the turbine blade, the cladding material in the tip repair area is ground and reshaped. This grinding and reshaping can be performed by mounting a grinding head on a CNC machine tool. Specifically, an 80-150# corundum wheel can be used for rough grinding, followed by a 300-800# diamond grinding head for fine grinding, ensuring that the surface roughness Ra of the turbine blade tip is ≤1.6μm after reshaping.

[0051] S70. Perform solution treatment and aging on the leaves.

[0052] The repaired turbine blades are placed in a vacuum furnace for heat treatment, and the blade structure and properties are restored through solution treatment and aging.

[0053] The turbine blades after solution treatment and aging can also be subjected to overall flaw detection. After confirming that there are no more defects such as cracks and holes in the repaired turbine blades, the turbine blades are transferred to the next post-processing process for further processing.

[0054] In some embodiments of the laser cladding repair method for tip damage of aero-engine turbine blades in this application, the method for performing material inspection heat treatment on the turbine blade in step S10 is as follows: the turbine blade is placed in a heat treatment furnace, and the furnace is heated at a rate of 20-50°C / min. After the temperature in the heat treatment furnace reaches 420-680°C, it is held at that temperature for 0.2-1.8 hours. After the holding period, the turbine blade is cooled to room temperature along with the furnace.

[0055] An aero-engine turbine blade that has undergone material inspection and heat treatment, such as Figure 2As shown, the turbine blade includes a tip 1, a blade body 2, a shroud 3, and a root 4. Within the tip 1 region, there exists a region 11 of a dissimilar material that visually differs from the base material within the tip 1 region. This indicates that the tip 1 of the turbine blade was previously repaired using a dissimilar repair material, and this material remains within the tip 1 region. To prevent crack formation during laser cladding repair, the dissimilar material region 11 within the tip 1 region needs to be removed, and the dissimilar repair material from the blade's base material needs to be eliminated.

[0056] In some embodiments of the laser cladding repair method for turbine blade tip damage in this application, step S20 involves flaw detection of the turbine blade using methods such as fluorescent testing, X-ray testing, eddy current testing, or penetrant testing. These methods can detect the number and distribution range of hidden cracks within the turbine blade tip region. Markings are made on the corresponding areas using tools such as markers, dyes, or electro-etching pens to facilitate the removal of crack distribution areas.

[0057] In some embodiments of the laser cladding repair method for turbine blade tip damage of aero-engines in this application, in step S30, a small grinding device is used to grind the area with crack markings in the blade tip 1 region to remove the matrix material with internal cracks, so that the remaining matrix material no longer has cracks inside.

[0058] If there is a foreign material area 11 in the blade tip area 1, use a small grinding device to grind the foreign material area 11 to remove the foreign repair material in the foreign material area 11, so that there is no foreign material in the remaining base material.

[0059] In a preferred embodiment of the laser cladding repair method for turbine blade tip damage of this application, after removing the cracked area inside the blade tip 1 region in step S30, an operation step S35 is further provided. In step S35, the turbine blade is subjected to flaw detection again using a flaw detection device to detect whether there are any remaining cracks that have not been completely removed in the blade tip 1 region.

[0060] If no cracks remain in the blade tip 1 area, continue with the subsequent steps; if cracks still exist in the blade tip 1 area, mark the cracked area and execute step S30 again, using a small grinding device to grind the marked cracked area, remove the cracked matrix material in the area, and ensure that no cracks remain in the matrix material.

[0061] In some embodiments of the laser cladding repair method for tip damage of aero-engine turbine blades in this application, such as Figure 3 and Figure 4As shown, in step S40, the thickness of the exhaust edge 12 at different positions within 1 cm below the removed end face after removing the cracked matrix material in the blade tip 1 region is measured, and the measurement results are compared with the original design thickness of the exhaust edge 12 to understand the aging and damage of the exhaust edge 12 during the operation of the aero-engine.

[0062] Using 80% of the original design thickness of the exhaust edge 12 as a comparison standard, if the measured thickness of the exhaust edge 12 is less than the comparison standard, a mark is made at the measurement point indicating that reinforcement is required, and the minimum thickness point 13 of the exhaust edge 12 is measured nearby. Then, the part of the exhaust edge 12 with a thickness less than the comparison standard is reinforced with the minimum thickness point 13 as the center.

[0063] If the thickness of each point on the measured exhaust edge 12 is greater than the thickness comparison standard, then there is no need to reinforce the exhaust edge 12. Proceed to step S50 to perform laser cladding repair on the removed end face of the blade tip 1.

[0064] In a preferred embodiment of the laser cladding repair method for turbine blade tip damage in this application, laser cladding is performed on the outer edge of the exhaust edge 12 to form a structure resembling... Figure 3 and Figure 4 The reinforcing rib 14 shown is used to reinforce the weak area of ​​the exhaust edge 12, improve the tensile strength of the weak area of ​​the exhaust edge 12, and prevent the exhaust edge 12 from being torn by the shrinkage stress of the cladding layer during the laser cladding repair process.

[0065] When laser cladding reinforcement is performed on the outer edge of the exhaust edge 12, the laser cladding power used is 80-200W, the diameter of the laser spot formed by laser irradiation is 0.2-1.5mm, and the moving speed of the laser spot on the exhaust edge 12 is 1-5mm / s. Nickel-based high-temperature alloy powder with the same composition as the substrate material of the blade tip 1 is used as the additive material for laser cladding, with a powder feed rate of 1-6g / min. A reinforcing rib 14 with a thickness of 0.5-0.9mm is formed on the outer side of the exhaust edge 12. The length of the reinforcing rib 14 is controlled to extend 2-5mm inward from the point of minimum thickness on the exhaust edge 12, ensuring that the reinforcing rib 14 at least covers the weak areas on both sides where the thickness is less than 80% of the original design thickness.

[0066] In some embodiments of the laser cladding repair method for turbine blade tip damage of this application, in step S50, before laser cladding the removal surface of the blade tip 1, the blade tip 1 area is preheated by heat treatment. Heat treatment preheating is typically performed using a medium-frequency induction heating device in the laser cladding processing area. The induction heating coil of the medium-frequency induction heating device surrounds the removal surface area of ​​the blade tip 1, and is positioned within an area approximately 5 mm below the removal surface of the blade tip 1.

[0067] Control the power of the medium-frequency induction heating device so that the temperature of the cladding area of ​​blade tip 1 increases at a rate of 20-50℃ / min. After the temperature of the cladding area of ​​blade tip 1 rises to 400-700℃, adjust the power of the medium-frequency induction heating device to maintain the temperature of the cladding area of ​​blade tip 1 within the range of 400-700℃ for heat preservation.

[0068] During the laser cladding repair of the removed end of blade tip 1, the medium-frequency induction heating equipment is kept in a warm state. After the laser cladding process of blade tip 1 is completed, it is kept warm for 2-10 minutes. Then, the temperature of the cladding area of ​​blade tip 1 is controlled to be cooled down to room temperature at a cooling rate of 10-30℃ / min.

[0069] Preheating the cladding area of ​​the blade tip 1 by heat treatment can reduce the thermal stress during laser cladding of the removal surface, reduce the risk of cracking of the base material of the blade tip 1 during the cladding process, and improve the microstructure of the cladding layer material, thereby improving the quality of turbine blade repair.

[0070] In a preferred embodiment of the laser cladding repair method for turbine blade tip damage of this application, in step S50, when performing laser cladding on the removed end of the blade tip 1, a coaxial machine vision system is used to generate the cladding path on the removed end surface of the blade tip 1, ensuring that the cladding material formed by laser cladding is stacked on the substrate material of the blade tip 1 in a set order to form a set shape, such as... Figure 3 The repair material structure 15 shown covers the original design shape of the blade tip 1 and is close to the original design shape of the blade tip 1. Thus, the blade tip 1 can be restored to its original design shape by reshaping the repair material structure 15 with less processing.

[0071] During laser cladding repair, an infrared thermometer is used to measure the temperature of the molten pool formed on the substrate material in real time, and the measured temperature value is transmitted to the controller. The controller adjusts the power of the cladding laser beam based on the real-time temperature of the molten pool using a specific algorithm. By adjusting the heat generated by the laser beam, the controller provides feedback regulation to stabilize the molten pool temperature within the range of 1800-2400℃, ensuring the stability of the cladding layer thickness and width formed in a single cladding process, and maintaining the shape and size of the additive structure formed by laser cladding.

[0072] During laser cladding, the laser power is adjusted within the range of 200-500W depending on the real-time temperature of the molten pool. The feed rate of the same nickel-based superalloy powder used in additive manufacturing is 2-8 g / min, the moving speed of the cladding laser beam spot on the substrate material is 1-6 mm / s, and the height of the cladding layer formed after the molten pool solidifies is 2.5-4.0 mm. By stacking the cladding material layer by layer along a certain path on the removed end face of blade tip 1, the structure of the substrate material removed from the end of blade tip 1 is repaired, restoring the original design morphology of blade tip 1.

[0073] As a specific embodiment of the laser cladding repair method for the tip damage of aero-engine turbine blades in this application, the blade tip 1 of the aero-engine turbine blade is laser cladding repaired using the net-shape repair system disclosed in Chinese Utility Model Patent No. CN204825049U. Argon gas with a purity of 99.9% is introduced into the environmental control box of the net-shape repair system, and the pressure inside the environmental control box is maintained between 1-2 MPa to promote the discharge of inclusions and gases inside the cladding layer formed by laser cladding, thereby improving the quality of the cladding layer.

[0074] Argon gas is circulated in the environmental control box via a gas circulation system. Oxygen and water vapor are filtered out of the argon gas using oxygen and water vapor filters, keeping the oxygen and water content in the argon gas below 20 ppm. This effectively avoids the oxidation of additive materials in a high-temperature environment during laser cladding repair, effectively solves the problem of oxidation inclusions within the cladding layer, and ensures the quality of laser cladding repair of the blade tip.

[0075] The laser cladding repair method for tip damage of aero-engine turbine blades described in this application is also applicable to the repair of damage to small and medium-sized high-pressure turbine blades on other equipment, especially turbine blades whose tips have undergone multiple repairs, and can effectively prevent the initiation of cracks during the repair process.

[0076] In the description of this invention, the terms "one embodiment," "specific embodiment," "preferred embodiment," etc., refer to specific features, structures, materials, or characteristics described in connection with that embodiment or example, which are included in at least one embodiment or example of the invention. In this invention, illustrative expressions of the above terms do not necessarily refer to the same embodiment or example. Furthermore, the specific features, structures, materials, or characteristics described may be combined in any suitable manner in one or more embodiments or examples.

[0077] The above are all preferred embodiments of this application, and are not intended to limit the scope of protection of this application. Therefore, all equivalent changes made in accordance with the structure, shape and principle of this application should be covered within the scope of protection of this application.

Claims

1. A laser cladding repair method for tip damage of aero-engine turbine blades, characterized in that, Includes the following steps: S10. Perform material inspection and heat treatment on the turbine blades to confirm whether there are any areas requiring repair with different materials in the blade tip area; S20. Perform flaw detection on the turbine blades and mark the location of cracks in the blade tip region; S30. Remove the cracked portion and foreign material from the blade tip region; S40. Measure the thickness of the exhaust edge within a set distance from the blade tip removal end face. If the thickness is less than the original design thickness setting value, then reinforce the exhaust edge. S50. Laser cladding is performed on the blade tip in a protective gas environment to restore the removed portion of the blade tip; S60. Grind and reshape the cladding area of ​​the blade tip to restore the original shape of the blade tip; S70. Perform solution treatment and aging treatment on the leaves.

2. The laser cladding repair method for tip damage of aero-engine turbine blades according to claim 1, characterized in that, In step S10, the material inspection heat treatment method is as follows: the turbine blade is heated to 420-680°C at a heating rate of 20-50°C / min, held at that temperature for 0.2-1.8 hours, and then cooled to room temperature in the furnace.

3. The laser cladding repair method for tip damage of aero-engine turbine blades according to claim 1, characterized in that, In step S20, the method for detecting flaws in the turbine blade includes fluorescent testing, X-ray testing, eddy current testing, or penetrant testing. The detection content includes the number of cracks and the distribution range of cracks in the blade tip region.

4. The laser cladding repair method for tip damage of aero-engine turbine blades according to claim 1, characterized in that, In step S30, a small grinding device is used to remove all cracked areas. If there are areas to be repaired with foreign materials, the foreign materials in the repair areas are also removed.

5. The laser cladding repair method for tip damage of aero-engine turbine blades according to claim 4, characterized in that, After step S30, the procedure further includes step S35: perform flaw detection on the turbine blade again. If there is a crack in the blade tip area, execute step S30 again.

6. The laser cladding repair method for tip damage of aero-engine turbine blades according to claim 1, characterized in that, In step S40, the thickness of the exhaust edge within 1 cm of the removed end face of the blade tip is measured. If the minimum value of the exhaust edge thickness is less than 80% of the original design thickness, the exhaust edge is reinforced.

7. The laser cladding repair method for tip damage of aero-engine turbine blades according to claim 6, characterized in that, The method for reinforcing the exhaust edge is to form a reinforcing rib by laser cladding on the outer edge of the exhaust edge. The reinforcing rib extends 2-5mm from the point of minimum thickness of the exhaust edge to both inner sides, and has a thickness of 0.5-0.9mm.

8. The laser cladding repair method for tip damage of aero-engine turbine blades according to claim 1, characterized in that, In step S50, before laser cladding the blade tip, the blade tip area is preheated by heat treatment. The heating rate of the heat treatment preheating is 20-50℃ / min, and the heating temperature is 400-700℃. After laser cladding the blade tip, the temperature is maintained for 2-10 minutes, and then cooled to room temperature at a cooling rate of 10-30℃ / min.

9. The laser cladding repair method for tip damage of aero-engine turbine blades according to claim 8, characterized in that, In step S50, when performing laser cladding on the blade tip, the cladding path of the blade tip is generated using coaxial vision, the temperature of the molten pool formed by laser cladding is measured in real time using an infrared thermometer, and the temperature of the molten pool is controlled to 1800-2400℃ by controlling the laser power. The laser power adjustment range is 200-500W, the powder feeding rate is 2-8g / min, the cladding speed is 1-6mm / s, and the height of the cladding layer formed after the molten pool solidifies is 2.5-4.0mm.

10. The laser cladding repair method for tip damage of aero-engine turbine blades according to claim 9, characterized in that, The protective gas environment is argon gas with a purity of 99.9% and a pressure of 1-2 MPa, and the water and oxygen content in the protective gas environment is below 100 ppm.

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