Bi-material turbine rotor blade and manufacturing method

Abstract

The invention provides a double-material turbine rotor blade and a manufacturing method. A target material is selected according to a heating analysis result of a part structure and material characteristics; classified slicing treatment is conducted based on the three-dimensional model of the part structure, a forming path scheme of electron beam selective melting of the double-powder-barrel system is planned, and layer-by-layer alternate deposition areas of the first type of materials and the second type of materials are determined; and alternative additive forming and interlayer sintering treatment are conducted according to the target material, and the double-material turbine rotor blade is obtained. Compared with traditional casting, overall improvement of the part structure performance is achieved based on the electron beam selective melting forming technology, the method is simple in process route, high in manufacturing efficiency and low in cost, and a technological innovation is provided for weight reduction, performance improvement, cost reduction, efficiency improvement and agile manufacturing of the aero-engine.

Description

Technical Field

[0001] The present invention belongs to the technical field of aero-engine manufacturing, and relates to a dual-material turbine rotor blade and a manufacturing method thereof. Background Art

[0002] With the continuous increase in the power performance requirements of aero-engines, the requirements for manufacturing technologies for engine weight reduction, the service environment of blades, and tissue performance have been synchronously improved. Engine weight reduction is mainly achieved in two aspects. On the one hand, the structural design is optimized, and on the other hand, a lighter material is used to replace the original material for the engine. At present, the turbine inlet temperature of aero-engines basically reaches 2040K, the temperature difference between the internal and external airflows of turbine air-cooled blades can be as high as 1300K, and lightweight materials generally have poor high-temperature resistance.

[0003] For aero-engine turbine rotor blades, the service environment is surrounded by high-temperature gas that is constantly changing, and they also bear huge centrifugal forces, gas forces, and vibration loads generated by high-speed rotation. After replacing the lightweight material, the blade body of the turbine rotor blade is very prone to problems such as ablation, burning through, and clogging of film holes by high-temperature gas, seriously affecting the overall service life of the turbine and even the aero-engine. Therefore, how to significantly reduce the weight of turbine rotor blades while ensuring their high-temperature resistance, and at the same time meet the requirements of low cost and short cycle for rapid manufacturing, will surely be the direction of future breakthroughs in manufacturing technology. Summary of the Invention

[0004] Aiming at the problems existing in the prior art, the present invention provides a dual-material turbine rotor blade and a manufacturing method thereof. Compared with traditional casting, the overall tissue performance of the part is improved, and this method has a simple process route, high manufacturing efficiency, and low cost, providing a technological innovation for aero-engines to achieve weight reduction, performance improvement, cost reduction, efficiency increase, and agile manufacturing.

[0005] The present invention is realized through the following technical solutions:

[0006] A manufacturing method for a dual-material turbine rotor blade includes:

[0007] Performing a thermal analysis on the part structure;

[0008] Selecting target materials according to the thermal analysis results of the part structure and material characteristics; the target materials include a first type of material and a second type of material;

[0009] Based on the three-dimensional model of the part structure, performing classification slicing processing, planning the forming path scheme of electron beam selective melting of a dual powder barrel system, and determining the layer-by-layer alternating deposition areas of the first type of material and the second type of material;

[0010] Based on the layer-by-layer alternating deposition area of the first type of material and the second type of material, alternating additive forming and interlayer sintering treatment are carried out according to the target material to obtain a dual-material turbine rotor blade.

[0011] Preferably, a thermal analysis is performed on the part structure, specifically:

[0012] A thermodynamic analysis is carried out on the position, temperature distribution, and heat flow direction of the turbine rotor blade under the action of high-temperature gas in the aeroengine.

[0013] Preferably, the alternating additive forming and interlayer sintering treatment is specifically as follows:

[0014] Use the first powder bucket and the left scraper to deposit the first type of material to form the first area of the current layer;

[0015] Increase the overall preheating power of the first area to achieve surface sintering and reach a semi-cured state;

[0016] Switch to the second powder bucket and the right scraper to deposit the second type of material to form the second area of the current layer;

[0017] Increase the preheating power of the second area to achieve surface sintering and reach a semi-cured state;

[0018] Repeat the above steps, and the deposition and interlayer bonding of the dual materials are completed alternately in areas until the overall forming of the dual-material turbine rotor blade.

[0019] Preferably, the first type of material is titanium aluminide alloy, and the second type of material is nickel-based superalloy.

[0020] Preferably, the nickel-based superalloy uses GH3536 superalloy material; the titanium aluminide alloy uses Ti4822 intermetallic compound material.

[0021] Preferably, the middle blade root and blade crown parts of the turbine rotor blade are formed with Ti4822 intermetallic compound material; the blade body part of the turbine rotor blade is formed with GH3536 superalloy material.

[0022] Preferably, the layer slicing process divides the functional area according to the heat gradient of the blade, including a low-temperature area and a high-temperature area.

[0023] Preferably, a transition area is provided between the alternating layers of the low-temperature area and the high-temperature area.

[0024] Preferably, the turbine rotor blade is subjected to vacuum heat treatment after final forming.

[0025] Preferably, a dual-material turbine rotor blade is obtained based on the manufacturing method of the above-mentioned dual-material turbine rotor blade.

[0026] Compared with the prior art, the present invention has the following beneficial technical effects:

[0027] A dual-material turbine rotor blade and manufacturing method of the present invention are applicable to the integrated manufacturing of dual-materials for turbine rotor blades of aero-engines. Specifically, based on the electron beam selective melting forming technology, for the turbine rotor blades of aero-engines, a dual-powder barrel form is adopted for layer-by-layer forming manufacturing to realize the integrated forming manufacturing of the blade body and other parts of the turbine rotor blade with dual materials. While ensuring the high-temperature resistance of the turbine rotor blade, a significant weight reduction of the part is achieved. Compared with traditional casting, an overall improvement in the tissue performance of the part is realized, and this method has a simple process route, high manufacturing efficiency, and low cost. This method can achieve a significant weight reduction of the turbine rotor blade while ensuring the high-temperature resistance of the turbine rotor blade of the aero-engine, providing a technological innovation for the aero-engine to achieve weight reduction, performance improvement, cost reduction and efficiency increase, and agile manufacturing.

[0028] Furthermore, for the electron beam selective melting forming technology of the present invention, compared with traditional investment casting, since the manufacturing environment of this invention method is a high-energy beam heat source manufacturing in a vacuum environment, the formed tissue has a higher density, fewer defects / inclusions, and better performance, effectively solving problems such as polycrystals, cracks, recrystallization, wall thickness deviation, and low casting qualification rate brought by casting technology, and providing a technological innovation for the quality improvement of turbine rotor blades of aero-engines.

[0029] Furthermore, for the turbine rotor blade body that is easily ablated and burned through by high-temperature gas, the GH3536 superalloy material is used for manufacturing, and the lighter Ti4822 intermetallic compound material is used for manufacturing other positions. Since the blade body uses the GH3536 material with better high-temperature resistance, the high-temperature resistance of the part is ensured. Since the Ti4822 material is used for the positions above the blade crown and below the middle blade root, a significant weight reduction of the overall turbine rotor blade is achieved, providing a technological innovation for the weight reduction and performance improvement of the aero-engine. Description of the Drawings

[0030] Figure 1 Schematic diagram of the heat analysis of the part structure of the embodiment;

[0031] Figure 2 Schematic diagram of the forming at the position below the middle blade root of the embodiment;

[0032] Figure 3 Schematic diagram of the forming at the blade body position of the embodiment;

[0033] Figure 4 Schematic diagram of the forming at the position above the blade crown of the embodiment;

[0034] Figure 5 Schematic diagram of the completed part manufacturing of the embodiment. Detailed Embodiments

[0035] The present invention will be further described in detail below in conjunction with specific embodiments, which are explanations of the present invention rather than limitations.

[0036] In order to enable those skilled in the art to better understand the technical solution of the present invention, the technical solution in the embodiments of the present invention will be clearly and completely described below 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 making creative efforts shall fall within the scope of protection of the present invention.

[0037] The method of the present invention belongs to the technical field of aeroengine manufacturing and is applicable to the dual-material integrated manufacturing of aeroengine turbine rotor blades. Specifically, based on the electron beam selective melting forming technology, for aeroengine turbine rotor blades, the Ti4822 intermetallic compound material and the GH3536 superalloy material are formed in layers in a dual powder barrel form, realizing the dual-material integrated forming manufacturing of the blade body and other parts of the turbine rotor blade. While ensuring the high temperature resistance of the turbine rotor blade, the weight of the part is significantly reduced. Compared with traditional casting, the overall tissue performance of the part is improved. Moreover, this method has a simple process route, high manufacturing efficiency, and low cost, providing a technological innovation for the aeroengine to achieve weight reduction, performance improvement, cost reduction, efficiency increase, and agile manufacturing. The inventive method is an integrated manufacturing method for high-performance, low-cost, dual-material aeroengine turbine rotor blades using the dual powder barrel electron beam selective melting forming manufacturing technology.

[0038] The invention belongs to the technical field of aeroengine manufacturing and relates to an integrated manufacturing method for high-performance, low-cost, dual-material aeroengine turbine rotor blades using the dual powder barrel electron beam selective melting forming manufacturing technology. The specific technical solution is realized through the following steps:

[0039] (1) Thermal analysis of the part structure: Analyze the position, temperature, and direction of the part under high-temperature gas in the aeroengine.

[0040] (2) Material selection and design: Select materials that can meet the service environment according to the results of the thermal analysis of the part structure and the material characteristics.

[0041] (3) Design of the forming path plan: Classify and slice the three-dimensional model according to the part structure to determine the forming path and plan of the electron beam selective melting of the dual powder barrel system.

[0042] (4) Forming of the first type of material: According to the forming plan of the part, first select the first type of material of the left scraper and the left powder barrel for additive forming.

[0043] (5) Intermediate layer sintering: Increase the preheating power of the powder layer to sinter and solidify the surface layer of the first type of material.

[0044] (6) Forming of the second type of material: According to the forming plan of the part, the second type of material of the right scraper and the right powder barrel is selected to continue the additive forming.

[0045] (7) Intermediate layer sintering: Increase the preheating power of the powder layer to sinter and solidify the surface layer of the second type of material.

[0046] (8) Forming of the first type of material: According to the forming plan of the part, continue to change back to the first type of material of the left scraper and the left powder barrel to continue additive forming.

[0047] (9) Manufacturing is completed layer by layer.

[0048] The following is a detailed description of steps (1)-(9):

[0049] Step (1) Thermal analysis of part structure

[0050] Thermodynamic modeling: Based on the three-dimensional geometric model of the aircraft engine turbine rotor blade, a transient thermodynamic simulation model is established through finite element analysis software (such as ANSYS, ABAQUS), and the gas temperature field parameters are input.

[0051] Determination of heat flow direction: According to the direction of gas flow under the engine working state, determine the heat flow impact area of the blade body, blade crown and blade root, and mark the location of maximum thermal stress concentration (such as the front 1 / 3 area of the blade body).

[0052] Temperature gradient division: The blade is divided into a high-temperature exposure zone (the area in direct contact with the gas), a medium-temperature transition zone (near the internal cooling channel), and a low-temperature load-bearing zone (the blade root and connecting structure) to generate a temperature gradient distribution map.

[0053] Step (2) Material selection and design

[0054] Material selection for high temperature areas: For high temperature exposure areas, choose nickel-based high temperature alloys (such as GH3536 high temperature alloy materials);

[0055] Material selection for low temperature zone: For low temperature load-bearing zone, choose lightweight titanium aluminum alloy (such as Ti4822 intermetallic compound);

[0056] Interface compatibility verification: Through thermal expansion coefficient matching analysis, the interface bonding stability of the two types of materials under thermal cycling is ensured.

[0057] Step (3) Forming path design

[0058] 3D model layering: Using an adaptive slicing algorithm, the blade model is layered along the Z axis (single layer thickness: 30-50 μm). According to the temperature gradient map in step (1), the deposition areas of the first type of material (titanium-aluminum intermetallic compound) and the second type of material (nickel-based high-temperature alloy) are divided in each layer.

[0059] Path planning: Set the alternating deposition order of the two powder buckets, specifically:

[0060] Layer 1: Titanium-aluminum intermetallic compound deposition (left powder barrel) → middle sintering → Titanium-aluminum intermetallic compound deposition (right powder barrel) → middle sintering;

[0061] Second layer: nickel-based high-temperature alloy deposition (right powder barrel) → middle sintering → nickel-based high-temperature alloy deposition (left powder barrel) → middle sintering;

[0062] Repeat alternately until completed.

[0063] Step (4) Forming the first type of material

[0064] Powder laying: The left scraper scrapes the first type of material (Ti4822 intermetallic compound material powder) from the left powder barrel to the forming substrate, with a powder laying thickness of 30μm and an accuracy of ±2μm;

[0065] Electron beam melting: using electron beam energy density of 80-100J / mm 3 , scanning speed 800-1200mm / s, melting the alloy area line by line;

[0066] Interlayer bonding control: The electron beam focus offset is set to ±0.05mm to eliminate unfused defects.

[0067] Step (5) Intermediate layer sintering

[0068] Local preheating strengthening: Increase the electron beam preheating power from the conventional value (500W) to 600-750W, and perform a second scan on the surface of the formed high-temperature alloy layer;

[0069] Sintering parameters: the scanning interval was expanded to 200 μm, and the scanning speed was reduced to 200 mm / s, so that the powder surface reached a semi-molten state (solidification rate ≥ 95%) and formed a densified transition layer.

[0070] Step (6) Forming the second type of material

[0071] Equipment switching: the right scraper switches to the right powder barrel to lay the second type of material (GH3536 high-temperature alloy material powder), and the powder laying synchronization accuracy is ≤0.01mm;

[0072] Differentiated melting: Adjust the electron beam parameters to an energy density of 60-80 J / mm 3, with a scanning speed of 1500 - 1800 mm / s, to avoid causing a heat affected zone (HAZ ≤ 50 μm) to the underlying superalloy.

[0073] Step (7) Intermediate layer sintering

[0074] Repeat step (5) to perform local sintering on the titanium aluminide layer, adjust the preheating power to 550 - 700 W, and shorten the sintering time by 10% - 15% to avoid excessive oxidation of the titanium aluminide.

[0075] Step (8) Forming of the first type of material

[0076] Switch back to the left powder bucket and left scraper, continue depositing the titanium aluminide according to the parameters in step (4), and simultaneously monitor the interlayer temperature in real-time through infrared temperature measurement (target value: 1050 ± 20 °C) to ensure interfacial metallurgical bonding.

[0077] Step (9) Completing the manufacturing layer by layer

[0078] Loop control: Repeat steps (4) - (8) until all layers are stacked;

[0079] Interface transition treatment: Between the alternating layers of nickel-based superalloy and titanium aluminide, use electron beam gradient energy scanning (energy linearly decreases from 100 J / mm 3 to 60 J / mm linearly 3 ) to form a 0.08 mm thick metallurgical transition zone;

[0080] Post-treatment: After forming, place the blade in a vacuum furnace, heat it to 800 - 1000 °C at a rate of 5 - 10 °C / min, hold for 2 - 4 hours, and then cool it in the furnace to below 200 °C to eliminate residual stress.

[0081] Ti4822 intermetallic compound material: Also known as TiAl4822, it is a γ-TiAl-based intermetallic compound and belongs to the second-generation cast TiAl alloy (typical composition is Ti - 48Al - 2Cr - 2Nb); it has excellent high-temperature strength and oxidation resistance, and the service temperature can reach 700 - 800 °C; the density is about 3.8 - 4.2 g / cm 3 , only about 1 / 2 of that of nickel-based superalloy, and it is an ideal lightweight high-temperature-resistant material in the aerospace field; the specific strength (strength / density) is significantly higher than that of traditional titanium alloys and superalloys; it is mainly used in the aerospace field, such as engine blades, structural components, etc., due to its high-temperature resistance and corrosion resistance characteristics.

[0082] The GH3536 superalloy material, also known as GH536, is a nickel-based superalloy with good high-temperature strength, oxidation resistance, and corrosion resistance. Its main components include nickel, chromium, molybdenum, etc. It is suitable for high-temperature environments, has good hot and cold workability and weldability, and can be made into complex components through processes such as forging, rolling, and welding. It is widely used in the fields of aerospace, energy, and chemical industry, such as aircraft engine combustion chambers, turbine blades, etc. Its high-temperature strength has a tensile strength of 650 MPa at 900 °C and can still maintain 400 MPa at 1000 °C. Its oxidation resistance shows an oxidation rate of only 0.001 g / cm 2 . Its corrosion resistance: It performs excellently in corrosive media containing sulfur, chlorine, etc., such as a corrosion rate of less than 0.02 mm / year in a sulfur-containing atmosphere at 650 °C.

[0083] This invention is mainly based on the electron beam selective melting forming technology. For the turbine rotor blade of an aircraft engine, a double powder bucket form is used to perform layer-by-layer forming manufacturing on the Ti4822 intermetallic compound material and the GH3536 superalloy material, realizing the integrated forming manufacturing of the double materials of the blade body and other parts of the turbine rotor blade. While ensuring the high-temperature resistance of the turbine rotor blade, it realizes a significant weight reduction of the part. With the technical advantages of electron beam selective melting forming, compared with traditional casting, it realizes an overall improvement in the tissue properties of the part, and this method has a simple process route, high manufacturing efficiency, and low cost. This invention method provides a technological innovation for the aircraft engine to achieve weight reduction, performance improvement, cost reduction and efficiency increase, and agile manufacturing.

[0084] Example 1

[0085] A manufacturing method for the turbine rotor blade of an aircraft engine. The steps are as follows:

[0086] (1) Thermal analysis of the turbine rotor blade structure: Analyze the position, temperature, and direction of the high-temperature gas on the turbine rotor blade in the aircraft engine, as shown in Figure 1 .

[0087] (2) Material selection and design: Select materials that can meet the service environment according to the results of the thermal analysis of the turbine rotor blade structure and material characteristics.

[0088] (3) Forming path scheme design: Classify and slice the three-dimensional model according to the turbine rotor blade structure to determine the forming path and scheme of the electron beam selective melting of the double powder bucket system.

[0089] (4) Forming of the Ti4822 intermetallic compound material at the position below the middle blade root: According to the forming scheme of the turbine rotor blade, first select the Ti4822 intermetallic compound material of the left scraper and the left powder bucket for additive forming, as shown in Figure 2 .

[0090] (5) Intermediate layer sintering: Increase the preheating power of the powder layer to sinter and solidify the surface layer of the Ti4822 intermetallic compound material.

[0091] (6) Forming of the GH3536 superalloy material at the blade body position: According to the forming scheme of the turbine rotor blade, continue with additive forming using the GH3536 superalloy material of the right scraper and the right powder bucket, as shown in Figure 3 .

[0092] (7) Intermediate layer sintering: Increase the preheating power of the powder layer to sinter and solidify the surface layer of the GH3536 superalloy material.

[0093] (8) Forming of the Ti4822 intermetallic compound material at the position above the blade crown: According to the forming scheme of the turbine rotor blade, continue to switch back to the Ti4822 intermetallic compound material of the left scraper and the left powder bucket for additive forming, as shown in Figure 4 .

[0094] (9) Complete the manufacturing layer by layer, as shown in Figure 5 .

[0095] After forming, the blade is placed in a vacuum furnace, heated to 950°C at a rate of 5°C / min, held for 3 hours, and then furnace-cooled to below 200°C to eliminate residual stress.

[0096] The inventive method innovatively proposes a manufacturing method for a dual-material turbine rotor blade. By leveraging the advantages of the dual-powder bucket electron beam selective melting forming technology, the inventive method selectively manufactures dual materials in a layered integrated part, which is an integrated manufacturing method for achieving significant weight reduction, performance improvement, and cost and cycle reduction of the turbine blade of an aeroengine. The main innovation points are as follows:

[0097] (1) By leveraging the advantages of the electron beam selective melting forming technology, the inventive method realizes the integrated manufacturing of dual materials at different layer positions of the same part of the turbine rotor blade of an aeroengine. For the blade body of the turbine rotor that is prone to being ablated and burned through by high-temperature gas, the GH3536 superalloy material is used for manufacturing, and the lighter Ti4822 intermetallic compound material is used for manufacturing at other positions. Since the blade body uses the GH3536 material with better high-temperature resistance, the high-temperature resistance of the part is ensured. Since the Ti4822 material is used at the positions above the blade crown and below the middle blade root, significant weight reduction of the overall turbine rotor blade is achieved, providing technological innovation for weight reduction and performance improvement of the aeroengine.

[0098] (2) By leveraging the advantages of electron beam selective melting forming technology, the inventive method realizes the forming and manufacturing of GH3536 and Ti4822 materials. Compared with traditional investment casting, since the manufacturing environment of the inventive method is high-energy beam heat source manufacturing in a vacuum environment, the prepared structure has higher density, fewer defects / inclusions, and better performance, effectively solving problems such as polycrystals, cracks, recrystallization, wall thickness deviation, and low casting qualification rate caused by casting technology, and providing technological innovation for improving the quality of aeroengine turbine rotor blades.

[0099] (3) By adopting the dual powder barrel electron beam selective melting forming technology, the inventive method realizes the integrated manufacturing of dual materials for turbine rotor blades. Compared with traditional investment casting technology, it saves the development of casting molds, has a simple process route, and high material utilization rate, which will effectively reduce the manufacturing cost of aeroengine parts, shorten the design scheme verification cycle, and realize the integrated, rapid, and low-cost manufacturing of complex structures and dual materials for aeroengine turbine rotor blades, providing technological innovation for the rapid trial production and verification of aeroengine.

[0100] In the description of the present invention, it should be understood that the orientation or positional relationship indicated by terms such as "center", "longitudinal", "transverse", etc. is based on the orientation or positional relationship shown in the drawings, and is only for the convenience of describing the present invention and simplifying the description, rather than indicating or implying that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation, and therefore cannot be understood as a limitation to the present invention.

[0101] It should be noted that the terms "first", "second", etc. in the description and claims of the present invention and the above drawings are used to distinguish similar objects, and do not necessarily need to describe a specific order or sequence. It should be understood that such data can be interchanged under appropriate circumstances, so that the embodiments of the present invention described here can be implemented in an order different from 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 comprising 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.

[0102] In the present invention, unless otherwise clearly defined and 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 horizontal height of the first feature is higher than that of 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 horizontal height of the first feature is less than that of the second feature.

[0103] 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.

[0104] 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 related listed items.

[0105] 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 the illustrations in 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 using the technical content disclosed above, such as slight modifications, embellishments and evolutions, are all 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 scope of protection of the technical solution of the present invention.

[0106] 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 the illustrations in 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 using the technical content disclosed above, such as slight modifications, embellishments and evolutions, are all 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 scope of protection of the technical solution of the present invention.

Claims

1. A method for manufacturing a dual-material turbine rotor blade, characterized in that, including, performing a thermal analysis on the part structure; selecting target materials according to the thermal analysis results of the part structure and material properties; the target materials include the first type of material and the second type of material; performing classification slicing processing based on the three-dimensional model of the part structure, planning the forming path scheme of electron beam selective melting of the dual powder barrel system, and determining the layer-by-layer alternating deposition areas of the first type of material and the second type of material; based on the layer-by-layer alternating deposition areas of the first type of material and the second type of material, performing alternating additive forming and interlayer sintering processing according to the target materials to obtain a dual-material turbine rotor blade.

2. The manufacturing method of a dual-material turbine rotor blade according to claim 1, wherein, Performing a thermal analysis on the part structure, specifically: Performing a thermodynamic analysis on the position, temperature distribution, and heat flow direction of the turbine rotor blade under the action of high-temperature gas in the aeroengine.

3. A method for manufacturing a dual-material turbine rotor blade according to claim 1, characterized in that, The alternating additive forming and interlayer sintering processing, the specific process is: Using the first powder barrel and the left scraper to deposit the first type of material to form the first area of the current layer; Increasing the overall preheating power of the first area to achieve surface sintering and reach a semi-cured state; Switching to the second powder barrel and the right scraper to deposit the second type of material to form the second area of the current layer; Increasing the preheating power of the second area to achieve surface sintering and reach a semi-cured state; Repeating the above steps, the areas are alternately completed with the deposition and interlayer bonding of the dual materials until the overall forming of the dual-material turbine rotor blade.

4. A method for manufacturing a dual-material turbine rotor blade according to claim 3, characterized in that, The first type of material is titanium aluminide, and the second type of material is nickel-based superalloy.

5. A method for manufacturing a dual-material turbine rotor blade according to claim 4, characterized in that, The nickel-based superalloy uses GH3536 superalloy material; the titanium aluminide uses Ti4822 intermetallic compound material.

6. The manufacturing method of a dual-material turbine rotor blade according to claim 5, characterized in that, The middle blade root and blade crown parts of the turbine rotor blade are formed with Ti4822 intermetallic compound material; the blade body part of the turbine rotor blade is formed with GH3536 superalloy material.

7. A method for manufacturing a dual-material turbine rotor blade according to claim 1, wherein, The layer slicing processing divides the functional areas according to the heat gradient of the blade, including a low-temperature area and a high-temperature area.

8. A method for manufacturing a dual-material turbine rotor blade according to claim 7, characterized in that, A transition area is provided between the alternating layers of the low-temperature area and the high-temperature area.

9. A method for manufacturing a dual-material turbine rotor blade according to claim 1, characterized in that, After final forming, the turbine rotor blade is subjected to vacuum heat treatment.

10. A dual-material turbine rotor blade, prepared by the manufacturing method of a dual-material turbine rotor blade according to any one of claims 1-9.

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