Aero-engine blade and multi-material integrated manufacturing method

Abstract

The invention belongs to the technical field of additive manufacturing, and relates to an aero-engine blade and a multi-material integrated manufacturing method, and the method comprises the steps: carrying out the bearing stress analysis of a part, and obtaining the position of a stress defect of the part; selecting a high-performance material to replace the stress defect position of the part for local replacement reinforcement; and performing simulation verification on the area for replacing the high-performance material, after the service performance requirement is met, designing a forming path scheme of laser cladding of the double powder feeding systems, based on the forming path scheme, starting the double powder feeding systems to perform composite material increase, and performing layer-by-layer superposition to complete part manufacturing and forming, so as to obtain the aero-engine blade. According to the method, the weight of the aero-engine part can be effectively controlled while the selective performance of the key risk position is improved, the defects of a post-treatment strengthening process technology are overcome, and the negative influence on the shape structure of the part is reduced; and a technical innovation is provided for the aero-engine to realize performance improvement, cost reduction, efficiency improvement and part service life prolonging.

Description

Technical Field

[0001] The present invention belongs to the technical field of additive manufacturing, and relates to a multi-material integrated manufacturing method for aero-engine blades. Background Art

[0002] With the continuous increase in the dynamic performance requirements of aero-engines, the service environment, tissue performance, and service life of blades have been synchronously improved for manufacturing technologies. Although after-treatment strengthening processes such as laser shock peening or shot peening can be used to achieve selective strengthening of weak positions of blade performance after the blank manufacturing of aero-engine parts, currently, there is generally a problem of high production cost of the strengthening process. Even for closed complex structures, there is a situation where strengthening cannot be achieved due to limited space dimensions. With the continuous increase in the structural complexity and performance requirements of aero-engine parts, traditional after-treatment strengthening technologies are gradually forming technical barriers. If high-performance materials are selected for overall manufacturing, it will undoubtedly greatly increase the production and manufacturing costs and affect the part weight. Therefore, how to achieve selective high-performance manufacturing of specific positions of complex-structured parts while meeting the rapid manufacturing requirements of low cost and short cycle will surely be the direction for future breakthroughs in manufacturing technologies. Summary of the Invention

[0003] Aiming at the problems existing in the prior art, the present invention provides a multi-material integrated manufacturing method for aero-engine blades. While selectively improving the performance of key risk positions, this method effectively controls the weight of aero-engine parts, breaks through the technical drawbacks of after-treatment strengthening processes, and reduces their negative impacts on the shape and structure of the parts themselves, providing a technological innovation for aero-engines to achieve performance improvement, cost reduction and efficiency increase, and part life extension.

[0004] The present invention is realized through the following technical solutions: A multi-material integrated manufacturing method for aero-engine blades, comprising: Performing a bearing force analysis on the part to obtain the positions of the force defects of the part; Selecting a high-performance material to replace and perform local replacement strengthening for the positions of the force defects of the part; Performing a simulation verification on the area where the high-performance material is replaced. After meeting the service performance requirements, designing a forming path plan for laser cladding with a dual powder feeding system; Based on the forming path plan, enabling the dual powder feeding system to perform composite additive manufacturing, and layer-by-layer stacking to complete the part manufacturing and forming to obtain aero-engine blades.

[0005] Preferably, performing a bearing force analysis on the part to obtain the positions of the force defects of the part, specifically: Simulate the stress state of the part in the actual working environment through simulation software, where the stress state includes the type of force and the force difference, identify the performance weak areas and easy failure positions of the part, and obtain the stress defect positions of the part; Preferably, the stress defect types of the part include insufficient strength and poor fatigue resistance.

[0006] Preferably, the raw material of the part is selected as TC4 titanium alloy material, its tensile strength is 1100MPa, and its yield strength is 1000MPa.

[0007] Preferably, the high-performance material is selected as TB18 titanium alloy material, its tensile strength is 1350MPa, and its yield strength is 1250MPa.

[0008] Preferably, if the verification result does not meet the service requirements, readjust the high-performance material and process parameters.

[0009] Preferably, design the forming path plan of the double powder feeding system laser cladding, specifically: perform layer slicing processing according to the three-dimensional model of the part, divide the model into continuous forming areas and local strengthening areas according to the geometric characteristics of the part, and dynamically allocate the collaborative operation time sequence of the double powder feeding system.

[0010] Preferably, based on the forming path plan, enable the double powder feeding system for composite additive manufacturing, and complete the part manufacturing and forming by layer-by-layer stacking. The specific process is as follows: Use the first powder feeding system to perform additive manufacturing of the matrix structure in the raw material area of the part; when cladding to the position of the stress defect of the part, switch to the second powder feeding system, and use high-performance materials for local strengthening additive forming. After the local processing is completed, switch back to the first powder feeding system to perform additive forming of the remaining parts in the raw material area of the part, and complete the part manufacturing and forming by layer-by-layer stacking to obtain an aeroengine blade with multi-material integrated manufacturing.

[0011] Preferably, the position of the stress defect of the aeroengine blade is the root position of the blade.

[0012] Preferably, an aeroengine blade is obtained by the multi-material integrated manufacturing method of the aeroengine blade.

[0013] Compared with the prior art, the present invention has the following beneficial technical effects: The method of the present invention proposes a multi-material integrated manufacturing method for aero-engine blades. By leveraging the advantages of multi-system powder feeding laser cladding technology, this invention method selectively manufactures integrated parts with multiple materials, which is an integrated manufacturing method for improving the performance, reducing the weight, and cutting costs and cycle time of aero-engine blades. By leveraging the advantages of multi-system powder feeding laser cladding technology, the present invention method realizes the multi-material integrated manufacturing of the same part at different positions of aero-engine blades. For positions where the part performance is weak and prone to failure, high-performance materials with similar material compositions and microstructures are used for laser cladding gradient forming to reduce the interfacial effect of dissimilar materials, achieve selective performance improvement at designated positions of the part, and improve the overall performance of aero-engine blades, providing technological innovation for weight reduction, performance improvement, and component life extension of aero-engine blades. By adopting multi-system powder feeding laser cladding technology, this invention method realizes the integrated manufacturing of parts with multiple materials, avoids the traditional post-treatment strengthening process for manufacturing parts, and will directly eliminate the negative impact of post-treatment strengthening technology on the shape and structure of the parts themselves, providing technological innovation for high-quality manufacturing of aero-engine blades.

[0014] Furthermore, through the integrated manufacturing of multiple materials, this invention method will effectively reduce costs. At the same time, due to the reduction of traditional manufacturing processes, it will effectively reduce the manufacturing cost of aero-engine parts, shorten the manufacturing cycle, realize the integrated, rapid, and low-cost manufacturing of closed complex structures and multi-material parts of aero-engines, and provide technological innovation for the rapid trial production and verification of aero-engines. Brief Description of the Drawings

[0015] Figure 1 It is a schematic diagram of the load-bearing force situation of the part; Figure 2 It is a schematic diagram of the first ordinary manufacturing stage; Figure 3 It is a schematic diagram of the strengthening manufacturing stage; Figure 4 It is a schematic diagram of the second ordinary manufacturing stage; Figure 5 It is a schematic diagram of the completion of part manufacturing. Detailed Description of the Preferred Embodiments

[0016] The following further elaborates on the present invention with specific embodiments, which are explanations rather than limitations of the present invention.

[0017] To enable those skilled in the art to better understand the solution of the present invention, the following will clearly and completely describe the technical solutions in the embodiments of the present invention in conjunction with the accompanying drawings in the embodiments of the present invention. Obviously, the described embodiments are only a part of the embodiments of the present invention, rather than all the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those of ordinary skill in the art without creative efforts shall fall within the protection scope of the present invention.

[0018] The present invention is applicable to the integrated strengthening manufacturing of high-performance, low-cost, and multi-material aero-engine parts. Specifically, a multi-system powder feeding laser cladding technology is adopted. High-performance materials are selected to replace the raw materials for strengthening manufacturing at the positions where the part performance is weak and prone to failure, so as to realize the integrated manufacturing of different materials for the same part of the aero-engine. Without significantly changing the weight of the part, the service performance and service life of the part are significantly improved. At the same time, since this manufacturing method avoids the post-treatment strengthening process, the production and manufacturing costs will be significantly reduced, and the manufacturing efficiency will be improved. Through the positive selective expression of the advantages of multi-materials and multi-properties, a technological innovation is provided for the integrated manufacturing of aero-engines to achieve performance improvement, cost reduction and efficiency increase, and part life extension. The inventive method is an integrated strengthening manufacturing method for high-performance, low-cost, and multi-material aero-engine parts using a multi-system powder feeding laser cladding technology.

[0019] The invention belongs to the field of additive manufacturing technology and relates to an integrated strengthening manufacturing method for high-performance, low-cost, and multi-material aero-engine parts using a multi-system powder feeding laser cladding technology. The specific technical solution is realized through the following steps: (1) Analysis of the load-bearing and stress conditions of the part: Through software simulation of the load-bearing and stress conditions of the part in the actual working environment, the positions where the part performance is weak and prone to failure are analyzed and obtained, as shown in Figure 1 .

[0020] (2) Material selection and design: On the basis of the original designed material, by analyzing the material composition and tissue performance of the material, high-performance material blades with similar material composition and organizational structure are selected to replace the weak positions for strengthening manufacturing, and the materials of the remaining parts remain unchanged.

[0021] (3) Verification of load-bearing and stress: Through software, the load-bearing and stress conditions of the weak positions after selecting high-performance materials are simulated again. After meeting the service performance requirements of the working environment, the design of the forming path scheme is started.

[0022] (4) Design of the forming path scheme: The three-dimensional model is classified and sliced according to the part structure to determine the forming path and scheme of the double powder feeding system laser cladding.

[0023] (5) General manufacturing stage: According to the forming scheme of the part, first select the original design material of System 1 for additive forming, as shown in Figure 2 .

[0024] (6) Reinforcement manufacturing stage: When the part is formed to the position with weak performance, replace it with the high-performance material of System 2 for additive forming, as shown in Figure 3 . Through the design of the forming path scheme, additive forming can also be carried out using a dual system at the same time.

[0025] (7) Complete the forming manufacturing by layer-by-layer stacking, as shown in Figure 5 .

[0026] The following is the specific detailed process: Step (1): Analysis of part load-bearing Specific process: Simulation modeling: Based on the three-dimensional model of the aero-engine part (such as a blade), use finite element analysis software (such as ANSYS, ABAQUS, etc.) to simulate the stress environment of the part under actual working conditions (such as high temperature, high pressure, centrifugal force, vibration load, etc.).

[0027] Load application: According to the engine operating parameters (speed, temperature distribution, air flow pressure, etc.), define the boundary conditions and load types (static / dynamic load, thermal-mechanical coupling load, etc.).

[0028] Identification of weak areas: Through stress nephograms, strain distributions, and fatigue life prediction results, determine the stress concentration areas in the part (such as blade roots, dovetail joint surfaces), high-temperature creep-sensitive areas, or positions prone to fatigue cracks ( Figure 1 the marked areas).

[0029] Output result: Generate an analysis report to clarify the key areas to be strengthened and the types of their performance defects (such as insufficient strength, poor fatigue resistance).

[0030] Step (2): Material selection design Specific process: Substrate analysis: Detect the composition and microstructure of the original design material (such as γ' phase distribution, grain size).

[0031] Screening of high-performance materials: Select high-performance materials with a composition gradient close to that of the substrate (such as the difference ≤ 10%) and a matching coefficient of thermal expansion to ensure interface compatibility.

[0032] Matching of tissue properties: Verify the matching degree of the grain size and precipitation phase distribution of the high-performance material with the matrix material through metallographic experiments (such as the grain size difference < 20%) to avoid failure caused by sudden changes in tissue at the interface. Local replacement solution: High-performance materials are used for the weak areas identified in step (1) (such as the leading edge of the blade), and the original design materials are retained for the remaining parts to control costs.

[0033] Step (3): Bearing force verification Specific process: Secondary simulation modeling: Input the properties of the replaced high-performance materials (elastic modulus, yield strength, thermal conductivity, etc.) into the simulation software, and reconstruct the part model including the locally strengthened area.

[0034] Performance verification: Simulate the stress distribution, fatigue life, and high-temperature creep behavior of the strengthened area to ensure that the maximum stress value is lower than the allowable stress of the material and the life is increased by ≥ 30%.

[0035] Iterative optimization: If the verification fails (such as stress concentration at the interface), adjust the selection range of the high-performance material or optimize the powder feeding process parameters (such as laser power, interlayer cooling time).

[0036] Confirm the solution: Output the final material layout and process window as the basis for the subsequent forming path design; Step (4): Forming path scheme design Specific process: Model layer slicing: Slice the three-dimensional part model along the Z-axis direction (layer thickness 50 - 200 μm), and divide the continuous forming area (substrate) and the locally strengthened area according to geometric features.

[0037] Path planning logic: Dual-system collaborative strategy: System 1 (raw material part) is responsible for the formation of the substrate structure, and System 2 (high-performance material) is responsible for local strengthening; the path design needs to avoid laser head interference and optimize the overlapping sequence.

[0038] Dynamic timing allocation: Control the start and stop timing of the dual powder feeding systems through numerical control codes (G-codes) to ensure that the material switching timing is synchronized with the cladding trajectory.

[0039] Process parameter matching: Set gradient parameter combinations of laser power (500 - 3000 W), scanning speed (5 - 20 mm / s), and powder feeding rate (2 - 10 g / min) for different areas to reduce the difference in the heat affected zone.

[0040] Step (5): General manufacturing stage Specific process: Start the formation of the substrate: Load the powder of the original design material using the first powder feeding system (System 1), and cladding the substrate structure layer by layer according to the preset path.

[0041] Process control: The diameter of the laser focused spot is 0.2 - 1.0 mm, and the temperature of the molten pool is controlled at 50 - 100 °C above the liquidus temperature of the material.

[0042] The surface morphology of the cladding layer is monitored in real time, and the process parameters are adjusted through closed-loop feedback to ensure that the matrix density is ≥99.5%.

[0043] Transition zone treatment: When approaching the 2 - 3 layers before the weak performance area, gradually reduce the powder feeding rate to reserve an interface transition gradient for material switching.

[0044] Step (6): Strengthening manufacturing stage Specific process: Material switching mechanism: Single system switching: When the cladding layer reaches the weak area, the numerical control system pauses system 1 and starts the second powder feeding system (system 2), using high-performance materials, and the switching time ≤0.5 s.

[0045] Dual system synchronization: For complex areas (such as special-shaped curved surfaces), the two powder feeding systems cooperate, and co-deposition of the two materials is achieved through powder flow focusing technology.

[0046] Interface strengthening control: Gradient powder feeding (gradual change in mixing ratio) is adopted at the material switching interface to reduce the crack tendency caused by sudden composition changes.

[0047] High-frequency oscillating laser scanning (frequency 200 - 500 Hz) is used to refine the interface grains and improve the bonding strength.

[0048] Real-time quality monitoring: The morphology of the molten pool is monitored by an infrared thermal imager and a high-speed camera, and the laser power and powder feeding amount are dynamically adjusted to ensure that there are no pores and unfused defects in the strengthening area.

[0049] Step (7): Complete forming by layer-by-layer stacking Specific process: Interlayer process adjustment: At the junction of the ordinary material layer and the high-performance material layer, gradually adjust the laser power (such as from 2000 W to 1800 W) and the scanning speed (such as from 10 mm / s to 12 mm / s) to relieve the accumulation of thermal stress.

[0050] After each layer of cladding, it is cooled with an inert gas (Ar or N2) to control the interlayer temperature gradient ≤50 °C / s.

[0051] Final forming: Repeat steps (5) - (6) until the overall forming of the part is completed, and the surface roughness Ra of the formed part is ≤12.5 μm.

[0052] Post-treatment: Stress relief annealing (such as 800 °C × 2 h) is carried out on the formed part to eliminate the residual stress, and finally an integrated part with both a high-strength matrix and local super-performance is obtained (Figure 5 ).

[0053] The method of the present invention aims to select high-performance materials for strengthening manufacturing at positions where the performance of parts is weak and prone to failure through the multi-system powder feeding laser cladding technology. It is a positive selective high-performance, low-cost, and integrated strengthening manufacturing method for complex part structures of multiple materials. While selectively improving the performance of key risk positions, this method can effectively control the weight of aero-engine parts, break through the drawbacks of post-treatment strengthening processes, reduce their negative impact on the shape and structure of the parts themselves, and provide a technological innovation for improving the performance, reducing costs and increasing efficiency, and extending the service life of aero-engine parts.

[0054] Example 1 A strengthening manufacturing method for a wide-chord blade of a certain aero-engine. The steps are as follows: (1) Analysis of the load-bearing and stress conditions of the part: By simulating the load-bearing and stress conditions of the part in the actual working environment through software, it is analyzed that the position where the performance of the part is weak and prone to failure is the root of the blade, as shown in Figure 1 .

[0055] (2) Material selection and design: Based on the analysis of the load-bearing and stress conditions, the original designed material for this part is TC4 titanium alloy material (tensile strength ~1100 MPa, yield strength ~1000 MPa). For the weak performance position, a TB18 high-performance titanium alloy material with similar material composition and excellent tissue performance (tensile strength ~1350 MPa, yield strength ~1250 MPa) is selected for strengthening manufacturing, and the materials of the remaining parts remain unchanged.

[0056] (3) Verification of load-bearing and stress: By simulating the load-bearing and stress conditions of the TB18 material position through software again, the forming path scheme design is started after meeting the load-bearing and stress working environment.

[0057] (4) Design of the forming path scheme: The three-dimensional model is classified and sliced according to the structure of the wide-chord blade to determine the forming path and scheme of the double-powder-feeding system laser cladding.

[0058] (5) Ordinary manufacturing stage: According to the forming scheme of the part, first select the TC4 material of system 1 to perform additive forming on the tenon of the wide-chord blade, as shown in Figure 2 .

[0059] (6) Strengthening manufacturing stage: When the blade is formed to the root of the blade, that is, the weak performance position, replace it with the TB18 material of system 2 to perform additive forming, as shown in Figure 3 .

[0060] (7) Ordinary manufacturing stage: Continue to replace it with the TC4 material of system 1 to perform additive forming on the middle and upper parts of the blade body of the wide-chord blade, as shown in Figure 4 .

[0061] (8) Complete the forming manufacturing by layer-by-layer stacking, as shown in Figure 5 .

[0062] The systematic integration of the above steps solves the problem that it is difficult to balance performance and cost in traditional single-material additive manufacturing, and is particularly suitable for the efficient strengthening manufacturing of key components such as aero-engine blades and turbine disks.

[0063] The method of the present invention realizes the multi-material integrated manufacturing of the same part at different positions of an aero-engine blade by leveraging the advantages of multi-system powder feeding laser cladding technology. For the positions where the part has weak performance and is prone to failure, laser cladding gradient forming is carried out using high-performance materials with similar material compositions and microstructures to reduce the interfacial effect of dissimilar materials, achieve the selective improvement of the performance at the specified positions of the part, and improve the overall performance of the aero-engine blade, providing technological innovation for weight reduction, performance improvement and component life extension of aero-engines.

[0064] It should be noted that the terms "first", "second", etc. in the description and claims of the present invention and the above-mentioned drawings are used to distinguish similar objects, and do not necessarily have to be used to describe a specific order or sequence. It should be understood that such used data can be interchanged under appropriate circumstances so that the embodiments of the present invention described here can be implemented in an order other than those illustrated or described here. In addition, the terms "comprising" and "having" and any variations thereof are intended to cover non-exclusive inclusion. For example, a process, method, system, product or device that includes a series of steps or units does not necessarily have to be limited to those steps or units clearly listed, but may include other steps or units not clearly listed or inherent to these processes, methods, products or devices.

[0065] In the present invention, unless otherwise clearly defined and limited, the terms "installed", "connected", "connected", "fixed", etc. should be understood in a broad sense. For example, it can be a fixed connection, a detachable connection, or integrated; it can be a mechanical connection, an electrical connection, or a communication connection; it can be directly connected, or indirectly connected through an intermediate medium, and can be the internal communication of two components or the interaction relationship between two components. For those of ordinary skill in the art, the specific meanings of the above terms in the present invention can be understood according to specific circumstances. When a component is said to be "fixed to" another component, it can be directly on the other component or there may also be an intermediate component. When a component is considered to be "connected" to another component, it can be directly connected to the other component or there may be an intermediate component at the same time. When a component is considered to be "disposed on" another component, it can be directly disposed on the other component or there may be an intermediate component at the same time.

[0066] In the present invention, unless otherwise clearly defined or limited, the first feature being "on" or "under" the second feature may include direct contact between the first and second features, or may include the first and second features not being in direct contact but in contact through additional features therebetween. Moreover, the first feature being "above", "over" and "on top of" the second feature includes the first feature being directly above and obliquely above the second feature, or merely indicating that the first feature has a higher horizontal height than the second feature. The first feature being "under", "beneath" and "underneath" the second feature includes the first feature being directly below and obliquely below the second feature, or merely indicating that the first feature has a lower horizontal height than the second feature.

[0067] It should be understood that when used in this specification and the appended claims, the terms "comprises" and "comprising" indicate the presence of the described features, integers, steps, operations, elements and / or components, but do not preclude the presence or addition of one or more other features, integers, steps, operations, elements, components and / or their groups.

[0068] Unless otherwise defined, all technical and scientific terms used herein have the same meaning as commonly understood by one of ordinary skill in the technical field to which this invention belongs. The terms used in the description of the present invention herein are for the purpose of describing specific embodiments only and are not intended to limit the present invention. The term "and / or" as used herein includes any and all combinations of one or more of the associated listed items.

[0069] The above are only the preferred embodiments of the present invention and do not impose any form of limitation on the present invention; any ordinary technician in the industry can smoothly implement the present invention according to what is shown in the accompanying drawings of the specification and the above description; however, any equivalent changes made by those skilled in the art within the scope of the technical solution of the present invention by making some modifications, refinements and evolutions using the technical content disclosed above are equivalent embodiments of the present invention; at the same time, any equivalent changes, modifications and evolutions made to the above embodiments based on the essential technology of the present invention still fall within the protection scope of the technical solution of the present invention.

[0070] The above are only the preferred embodiments of the present invention and do not impose any form of limitation on the present invention; any ordinary technician in the industry can smoothly implement the present invention according to what is shown in the accompanying drawings of the specification and the above description; however, any equivalent changes made by those skilled in the art within the scope of the technical solution of the present invention by making some modifications, refinements and evolutions using the technical content disclosed above are equivalent embodiments of the present invention; at the same time, any equivalent changes, modifications and evolutions made to the above embodiments based on the essential technology of the present invention still fall within the protection scope of the technical solution of the present invention.

Claims

1. A multi-material integrated manufacturing method for aero-engine blades, characterized in that, including, conduct a bearing force analysis on the part to obtain the position of the force defect of the part; select a high-performance material to replace and locally strengthen the part according to the position of the force defect of the part; perform a simulation verification on the area where the high-performance material is replaced. After meeting the service performance requirements, design a forming path plan for laser cladding with a dual powder feeding system, Based on the forming path plan, enable the dual powder feeding system for composite additive manufacturing, and layer by layer stack to complete the manufacturing and forming of the part, obtaining an aero-engine blade.

2. The multi-material integrated manufacturing method of an aero-engine blade according to claim 1, characterized in that, Conduct a bearing force analysis on the part to obtain the position of the force defect of the part, specifically: Simulate the stress state of the part in the actual working environment through simulation software. Among them, the stress state includes the type of force and the difference in force, identify the performance weak areas and easy failure positions of the part, and obtain the position of the force defect of the part.

3. The multi-material integrated manufacturing method of an aero-engine blade according to claim 2, characterized in that, The types of force defects of the part include insufficient strength and poor fatigue resistance.

4. A multi-material integrated manufacturing method for an aero-engine blade according to claim 1, characterized in that, The raw material of the part is selected as TC4 titanium alloy material, with a tensile strength of 1100 MPa and a yield strength of 1000 MPa.

5. A multi-material integrated manufacturing method for an aero-engine blade according to claim 1, characterized in that The high-performance material is selected as TB18 titanium alloy material, with a tensile strength of 1350 MPa and a yield strength of 1250 MPa.

6. The multi-material integrated manufacturing method of an aero-engine blade according to claim 1, characterized in that, If the verification result does not meet the service requirements, readjust the high-performance material and process parameters.

7. A multi-material integrated manufacturing method for an aero-engine blade according to claim 1, characterized in that, Design a forming path plan for laser cladding with a dual powder feeding system, specifically: perform layer slicing processing according to the three-dimensional model of the part, divide the model into continuous forming areas and local strengthening areas according to the geometric characteristics of the part, and dynamically allocate the collaborative operation time sequence of the dual powder feeding system.

8. A multi-material integrated manufacturing method for an aero-engine blade according to claim 1, characterized in that, Based on the forming path plan, enable the dual powder feeding system for composite additive manufacturing, and layer by layer stack to complete the manufacturing and forming of the part. The specific process is: Use the first powder feeding system to perform additive manufacturing of the matrix structure in the raw material area of the part; when cladding to the position of the force defect of the part, switch to the second powder feeding system and use high-performance materials for local strengthening additive forming. After the local processing is completed, switch back to the first powder feeding system to perform additive forming on the remaining parts of the raw material area of the part, and layer by layer stack to complete the manufacturing and forming of the part, obtaining an aero-engine blade with multi-material integrated manufacturing.

9. The multi-material integrated manufacturing method of an aero-engine blade according to claim 1, characterized in that The position of the force defect of the aero-engine blade is the root position of the blade.

10. An aero-engine blade is obtained by the multi-material integrated manufacturing method of an aero-engine blade according to any one of claims 1-9.

Images