High-temperature-resistant aero-engine blade and preparation method thereof

Abstract

The application belongs to the technical field of aero-engine materials, and particularly relates to a high-temperature-resistant aero-engine blade and a preparation method thereof. The blade adopts titanium-aluminum alloy as a matrix, introduces two new types of materials of a core-shell structure self-healing reinforcing phase and a rare earth niobate iron elastic protective compound, and significantly improves the comprehensive performance of the blade in a high-temperature environment. The core-shell reinforcing phase takes titanium diboride as a core and boron nitride as a shell, and can realize self-healing function at high temperature. The rare earth niobate compound is modified by doping tantalum and rhenium, and has ferroelastic behavior and ultralow thermal conductivity. The preparation process comprises the following steps: mixing titanium-aluminum alloy powder, the reinforcing phase, silicon carbide short fibers and a binder, and then performing discharge plasma sintering to form; after mechanical processing, a protective coating is applied by plasma spraying; and finally, heat treatment is performed. The method simplifies the process, the prepared blade has excellent high-temperature resistance, oxidation resistance and mechanical properties, and is suitable for aero-engine hot end components.

Description

TECHNICAL FIELD

[0001] The application belongs to the technical field of aero-engine materials, and particularly relates to a high-temperature-resistant aero-engine blade and a preparation method thereof. BACKGROUND

[0002] As a cutting-edge product in the field of modern industrial technology, the performance level of an aero-engine directly determines the comprehensive efficiency and development prospects of an aircraft. Among the many core components of an engine, the performance of blades, especially turbine blades and compressor blades in a high-temperature gas environment, plays a decisive role in the overall thrust, fuel efficiency and working reliability of the engine. With the continuous development of aviation technology towards higher thrust-to-weight ratio and stronger thermal efficiency, the turbine inlet temperature of the engine continues to rise, which poses almost stringent requirements for the high-temperature resistance, long-term structural stability and environmental erosion resistance of blade materials. Although the nickel-based high-temperature alloy widely used in blade manufacturing has good high-temperature strength and toughness, its high density is not conducive to the lightweight design of the engine, and its upper limit of working temperature has gradually become difficult to meet the needs of the new generation of engines. Although titanium-based alloys perform excellently in lightweight design, they are prone to softening and oxide scale peeling in an oxidizing high-temperature environment of several hundred degrees Celsius, losing structural integrity. Therefore, the scientific research and engineering communities have turned their attention to new material systems such as ceramic matrix composites, which have the characteristics of high melting point, low density, excellent oxidation resistance and non-brittle fracture of composite materials, and are considered as a key direction to break through the bottleneck of existing technology. However, even these advanced materials still face severe challenges in actual application, such as the initiation of micro-cracks in the material under the coexistence of high temperature, high stress and water-oxygen medium, which becomes a channel for rapid inward diffusion of oxygen, eventually leading to oxidation embrittlement of the fiber reinforcement and matrix interface, and a sharp decline in material performance. This series of inherent problems constitutes the main technical obstacle restricting the development of high-performance aero-engine blades.

[0003] To address the above challenges, domestic and foreign researchers have carried out a lot of exploratory work. In terms of improving the protection ability of the material itself, early research focused on introducing compounds with self-healing ability, hoping that they would generate a flowable glass phase at high temperature to fill cracks and block the oxygen diffusion path. However, when such traditional self-healing materials are exposed to high-temperature steam environments, the protective glass phase they generate will volatilize and lose due to chemical reaction with water vapor, resulting in rapid decay of protection effect and insufficient durability. In terms of thermal protection, widely used thermal barrier coating materials will undergo harmful crystal structure transformation and densification sintering when exceeding a certain temperature threshold, not only causing a significant decline in their thermal insulation performance, but also peeling off due to increased internal stress, exposing the substrate material directly to high-temperature combustion gas. In recent years, some new material systems have shown potential application value, such as composite oxide coatings with ferroelastic behavior, whose unique micro-domain switching mechanism helps to disperse stress and delay crack propagation; and attempts to stabilize crystal structure and optimize thermal physical properties through specific element doping. However, most existing technical solutions are still limited to the improvement of a single performance, either focusing on substrate strengthening or surface protection, and have not formed a complete and coordinated technical system that enhances from the inside of the material to the outside protection. At the same time, many improvement schemes are accompanied by the complication of the preparation process and a significant increase in cost, making it difficult to meet the actual needs of large-scale and high-consistency manufacturing of aircraft engine blades.

[0004] In summary, the core problem in the current technical field of aircraft engine blade materials is the lack of a comprehensive material solution that can balance the high-temperature strength of the substrate, internal self-healing and oxidation resistance, and surface efficient thermal barrier protection, and its preparation process should be simplified as much as possible to have engineering application prospects. The durability, stability and reliability of existing material systems in extreme environments are still insufficient, and they cannot fully meet the pursuit of performance limits by future aircraft engines. Therefore, the industry is eagerly awaiting the emergence of a new material design concept that can achieve multi-level collaborative protection and be prepared through a relatively simplified process to break through the existing technical bottlenecks and provide key material support for the development of the next generation of aircraft engines. This is not only a theoretical problem of material science, but also a major engineering and technical issue related to the level of high-end equipment manufacturing and core competitiveness of the country. SUMMARY

[0005] In view of the deficiencies of the prior art, the purpose of the present application is to provide a high-temperature-resistant aircraft engine blade and a preparation method thereof.

[0006] In a first aspect of the present application, a preparation method of a high-temperature-resistant aircraft engine blade is provided, comprising the following steps:

[0007] S1, mix TiAl alloy powder, core-shell structure self-healing reinforcing phase, SiC short fiber and polycarbosilane solution in a mixer to obtain a mixture; load the mixture into a graphite mold, and adopt spark plasma sintering to form under the condition of temperature 1150-1250 DEG C, and keep for a certain time to obtain a blade blank;

[0008] S2, precisely process the blade blank by a five-axis numerical control machine tool to form a blade geometric shape; mix rare earth niobate iron elastic protective compound powder with HfO2-Y2O3 and polymethyl methacrylate microspheres, and spray on the surface of the blade by a plasma spraying device to form a blade with a protective coating;

[0009] S3, heat treat the blade with the protective coating under the protection of argon at 790-810 DEG C, and cool down with the furnace.

[0010] In the application, the preparation mechanism of the high-temperature-resistant aero-engine blade is reflected in the integration and synergistic effect of the multi-level material system. The blade substrate is densified by spark plasma sintering technology: the titanium aluminum alloy powder flows plastically under the action of high temperature and pressure, and the polycarbosilane binder pyrolyzes to form a silicon carbide ceramic phase. This process not only fills part of the pores, but also strengthens the interface bonding between the titanium aluminum matrix and the reinforcing phase through in-situ reaction. The core-shell reinforcing phase and the silicon carbide short fiber form a three-dimensional network in the matrix, in which the silicon carbide fiber improves the toughness of the material through the load transfer mechanism, and the core-shell particles play a role through the "crack pinning-healing" dual mechanism. When the matrix produces a crack, the crack tip deflects and branches when it encounters the core-shell particles, consuming the fracture energy; at the same time, the particle shell layer is oxidized to form a glass phase, realizing crack self-sealing. In the coating preparation stage, the rare earth niobate compound and the hafnium oxide-yttrium oxide mixture are deposited on the surface of the blade by plasma spraying, and the polymethyl methacrylate microspheres are decomposed to form a uniform distribution of microporous structure. These pores not only effectively reduce the thermal conductivity of the coating, but also provide strain tolerance for the coating to avoid the accumulation of thermal stress. The final heat treatment process promotes the interface interdiffusion between the coating and the matrix, forms a metallurgical bond, and releases internal stress through atomic rearrangement, and optimizes the organizational structure. The whole system realizes the long-term service of the blade in extreme environment through the multi-level synergies of matrix bearing, reinforcing phase toughening, self-healing protection and thermal barrier coating heat insulation.

[0011] As a preferred technical scheme of the application, in step S2, the thickness of the protective coating of the blade with the protective coating is 150-300 μm.

[0012] As a preferred technical scheme of the application, in step S3, the heat treatment time is 2-4 h.

[0013] As a preferred technical scheme of the present application, the preparation method of the core-shell structure self-healing reinforcing phase comprises: A1, mixing Ti powder and B powder to obtain a mixture, and molding the mixture to obtain a green body; then igniting the green body under argon protection to obtain a TiB2 porous preform; A2, placing the TiB2 preform in a CVD reactor, and reacting at 1100-1300 DEG C under the condition of BCl3 and NH3 as precursors.

[0014] In the present application, the design core of the core-shell structure self-healing reinforcing phase is to realize the crack self-repairing function of the material in a high temperature environment through a multi-stage reaction. The preparation process first mixes titanium powder and boron powder according to a specific mass ratio by using a self-propagating high-temperature synthesis technology, molds the mixture to obtain a green body, ignites the green body, and generates a porous preform with titanium diboride as the core under the condition of a violent exothermic reaction under argon protection. The reaction utilizes the mutual diffusion and bonding of titanium and boron atoms at high temperature to form a ceramic phase with high melting point and high hardness, and the porous characteristics of the ceramic phase provide a structural basis for subsequent shell coating. Subsequently, by using a chemical vapor deposition process, boron chloride and ammonia gas are used as precursors to make gas molecules penetrate into the pores of the preform and react under high temperature conditions to generate a boron nitride shell layer in situ on the surface of the titanium diboride core. This process relies on a gas-solid phase interface reaction, and by controlling the temperature and gas concentration, the boron nitride uniformly covers the surface of the core phase in the form of a dense film to form a complete core-shell structure. When micro-cracks occur in the composite material during service, the boron nitride shell layer will preferentially react with oxygen in the environment to generate a boron-oxygen glass phase. The molten glass phase rapidly fills the cracks under the driving force of capillary force to form a dense barrier layer, thereby blocking the further diffusion of oxygen inward, and realizing the self-healing function of the material. This mechanism effectively solves the problem of performance degradation caused by crack propagation of traditional materials in a high temperature water-oxygen environment.

[0015] As a preferred technical scheme of the present application, in step A1, the mass ratio of Ti powder to B powder is 69: (32-34).

[0016] As a preferred technical scheme of the present application, in step A2, the reaction time under the condition of 1100-1300 DEG C is 2-4h.

[0017] As a preferred technical scheme of the present application, the preparation method of the rare earth niobate iron elastic protective compound comprises: B1, dissolving Nd2O3, Nb2O5, Ta2O5 and ReO2 in dilute nitric acid to form a mixed solution; adding oxalic acid solution to generate an oxalate precipitate, filtering and washing with deionized water; dispersing the oxalate precipitate in deionized water again, adding polyethylene glycol, and spray drying to obtain a precursor powder; B2, sintering in an air atmosphere at 1450-1550 DEG C.

[0018] In the present application, the synthesis of rare earth niobate ferroelastic protective compounds is based on element doping and solid phase reaction to control the crystal structure, so as to realize the characteristics of ferroelasticity and ultra-low thermal conductivity. In the early stage of preparation, oxides of neodymium, niobium, tantalum and rhenium are mixed at the molecular level in the solution by oxalate co-precipitation method, and uniform precipitate is formed through the coordination reaction of oxalate ions and metal cations. This process ensures the uniform distribution of elements at the atomic scale, laying a foundation for the subsequent formation of homogeneous composite oxides. After spray drying of the precipitate, a series of solid phase reactions occur in the high temperature sintering stage: first, neodymium and niobium oxides react to form neodymium niobium oxide matrix phase, then tantalum atoms partially replace niobium crystal sites, and rhenium atoms partially replace neodymium crystal sites. This doping not only stabilizes the coexistence of fluorite and monoclinic crystal structures, but also introduces lattice distortion and defect dipoles, thereby inducing ferroelasticity. Under the action of stress, the ferroelastic domain of the compound will undergo reversible switching, absorbing strain energy through domain wall motion, promoting crack tip deflection and bridging, and significantly improving the fracture toughness of the coating. At the same time, the doping of tantalum and rhenium significantly enhances the phonon scattering effect, reducing the lattice thermal conductivity and making the material have excellent heat insulation performance. The final composite phase can maintain structural stability at high temperatures and adapt to the thermal expansion of the substrate through its ferroelastic behavior, avoiding coating peeling in thermal cycling.

[0019] As a preferred technical scheme of the present application, in step B1, the mass ratio of Nd2O3, Nb2O5, Ta2O5 and ReO2 is (70.0-72.0):(53.5-55.0):(12.5-13.5):(12.0-13.5).

[0020] As a preferred technical scheme of the present application, in step B2, the sintering time in air atmosphere at 1450-1550℃ is 5-8h.

[0021] In the second aspect of the present application, a high-temperature-resistant aero-engine blade prepared by the preparation method of the high-temperature-resistant aero-engine blade is provided, which comprises the following raw materials in parts by weight: TiAl alloy 60-75 parts; core-shell structure self-healing reinforcing phase 15-25 parts; rare earth niobate ferroelastic protective compound 10-20 parts; SiC short fiber 5-15 parts; HfO2-Y2O3 3-8 parts; polycarbosilane 2-5 parts; polymethyl methacrylate microspheres 1-3 parts.

[0022] Compared with the prior art, the present application has the following beneficial effects:

[0023] (1) The high-temperature-resistant aero-engine blade and the preparation method thereof provided by the application realize the significant improvement of the comprehensive performance of the blade in an extremely high-temperature environment through innovative material system design and optimized preparation process. Compared with the traditional technical solution, the most prominent technical effect of the application lies in the successful construction of a multi-level collaborative protection system from matrix strengthening, internal self-healing to surface protection. The system takes titanium-aluminum alloy as the matrix, introduces self-healing reinforcing phases with a special core-shell structure and rare earth niobate iron elastic protective compounds modified by element doping, and realizes the collaborative breakthrough of the material in high-temperature strength, oxidation resistance and thermal barrier performance. Specifically, the core-shell reinforcing phase can trigger a self-healing mechanism at high temperature, effectively filling the micro-cracks generated in the material, and fundamentally blocking the diffusion channel of oxygen; and the surface iron elastic protective layer has both ultra-low thermal conductivity and excellent thermal shock resistance due to its unique microstructure, thereby providing long-term and stable thermal protection for the matrix. This inside-out repair design concept enables the finally prepared aero-engine blade to maintain the structural integrity and stability of the performance under long-term high-temperature service conditions, and solves the technical bottleneck of the existing material system that the single performance is outstanding but the comprehensive performance is insufficient.

[0024] (2) In terms of specific performance indicators, the application exhibits comprehensive and excellent technical effects. First, the significant breakthrough in high-temperature resistance, the blade can work stably in a high-temperature environment far beyond the limit of traditional nickel-based alloys, the maximum use temperature is greatly improved, and the material density is significantly reduced, laying a solid foundation for the lightweight design and efficiency improvement of the aero-engine. In terms of oxidation resistance and self-healing performance, thanks to the phase change behavior of the shell material in the core-shell structure at high temperature, the generated flow glass phase can quickly and effectively heal the cracks, and the protection effect will not decay quickly even in a high-temperature gas environment rich in water vapor, showing much better durability than traditional self-healing materials. In terms of mechanical properties, the reasonable distribution of the reinforcing phase forms a stable strengthening network with the matrix, so that the blade material can still maintain high strength tensile strength and creep resistance at high temperature, and its high-temperature endurance life is several times that of traditional materials. At the same time, the introduction of the iron elastic compound coating dissipates energy through its unique domain switching mechanism, significantly improves the fracture toughness of the material, and makes it less likely to cause catastrophic damage when subjected to severe thermal shock. The collaborative improvement of these performance indicators constitutes the core competitiveness of the application in dealing with extreme working conditions of the aero-engine.

[0025] (3) The technical effects brought by the present application also lie in the simplification of the preparation process and the convenience of engineering application. By adopting the spark plasma sintering technology, rapid densification forming of the matrix material is realized, which greatly shortens the process flow and reduces the energy consumption; and the application of atmospheric plasma spraying technology makes the preparation of the surface protective coating of the complex curved blade efficient and uniform. The whole production process is smooth, avoiding the quality control difficulties common in traditional complex processes, which is conducive to realizing the consistency and stability of product performance, meeting the needs of large-scale production of aviation parts. In summary, the present application not only brings a qualitative leap in blade performance through the original innovation of the material system, but also ensures the feasibility and economy of the technical scheme through the integration and simplification of the process, providing key material technology support for the independent research and development of China's new generation of high-performance aero-engines, and has great theoretical research value and engineering application prospect. DETAILED DESCRIPTION

[0026] In order to facilitate the understanding of the present application, the present application lists the following examples. Those skilled in the art should understand that the examples are only to help understand the present application, and should not be regarded as specific limitations on the present application.

[0027] Some components in the examples and comparative examples are as follows:

[0028] The TiAl alloy powder is purchased from Western Metal Materials Co., Ltd.

[0029] The SiC short fiber is purchased from Ningbo Folong Advanced Materials Co., Ltd.

[0030] The polycarbosilane is purchased from Fudian Yayachemical Co., Ltd.

[0031] The polymethyl methacrylate microspheres are purchased from Guangzhou Tianyuan Silica Gel Technology Co., Ltd.

[0032] The five-axis numerical control machine tool is purchased from Beijing Jingdiao Technology Group Co., Ltd.

[0033] The HfO2-Y2O3 is purchased from Shanghai Yaohua Nanometer Material Co., Ltd.

[0034] The Ti powder is purchased from Beijing Antai Technology Co., Ltd.

[0035] The B powder is purchased from Yingkou Defeng New Material Technology Co., Ltd.

[0036] The BCl3 is purchased from Zhonghao Guangming Chemical Research and Design Institute Co., Ltd.

[0037] The NH3 is purchased from Hangzhou Oxygen Machine Group Co., Ltd.

[0038] The Nd2O3 is purchased from Qiandong Rare Earth Group Co., Ltd.

[0039] The Nb2O5 was purchased from Dongfang Tantalum Industry Co., Ltd.

[0040] The Ta2O5 was purchased from Hunan Jinwang Bismuth Industry Co., Ltd.

[0041] The ReO2 was purchased from Chengdu Zhongjian Material Optoelectronic Co., Ltd.

[0042] The oxalic acid was purchased from Shandong Yingxuan Industry Co., Ltd.

[0043] The polyethylene glycol was purchased from PEG-400 type of Jiangsu Sixin Technology Application Research Institute Co., Ltd.

[0044] The ordinary TiB2 powder was purchased from Qinghe County Zhongzhou Alloy Material Co., Ltd.

[0045] The ordinary yttrium stabilized zirconia powder was purchased from Beijing Huawei Ruikexin Chemical Technology Co., Ltd.

[0046] Example One:

[0047] The preparation of the core-shell structure self-healing reinforcing phase first carried out the synthesis of TiB2 porous preform, specifically using 69 g of titanium powder and 32 g of boron powder, the purity of the two powders was 99.9%, the particle size of the titanium powder was controlled at 1-3 μm, the particle size of the boron powder was controlled at 0.5-1 μm, in the three-dimensional mixer under the protection of argon, mixed at a speed of 200 r / min for 4 h to ensure uniformity, the mixed powder was loaded into the mold, and the green body was obtained under the pressure of 50 MPa; the green body was loaded into a special graphite crucible, and the reaction was ignited at 1300°C in a self-propagating high-temperature synthesis device to obtain a TiB2 porous preform with an ideal porosity of 45%; then the chemical vapor deposition of BN shell layer was carried out, the preform was placed in a CVD reactor, high-purity BCl3 and NH3 were used as precursors, Ar was used as carrier gas, the BCl3 flow rate was accurately controlled at 50 sccm, the NH3 flow rate was 150 sccm, the reaction temperature was 1180°C, the system pressure was 30 Pa, and the reaction was continued for 3 h, finally a dense BN shell layer with a thickness of 150 nm was formed on the surface of the TiB2 core.

[0048] The rare earth niobate iron elastic protective compound was prepared by oxalate coprecipitation method. 71.0 g of Nd2O3, 54.0 g of Nb2O5, 13.0 g of Ta2O5 and 12.5 g of ReO2 were accurately weighed, all of which had a purity of 99.99%, dissolved in a 2 mol / L dilute nitric acid solution, stirred at 80°C until completely clear, and then an excess of 20% 0.5 mol / L oxalic acid solution was added to form a uniform coprecipitate. After aging for 12 h, the precipitate was filtered and washed repeatedly with deionized water until the pH of the filtrate was 6. The obtained precipitate was redispersed and 5 g of polyethylene glycol-400 was added as a dispersant. The precursor powder was obtained by a spray dryer at an inlet temperature of 220°C and an outlet temperature of 100°C. Then the precursor powder was placed in a high-temperature sintering furnace, and the temperature was raised to 1500°C at a rate of 5°C / min, and sintered in air for 6 h to form a Ta and Re doped Nd3NbO7-NdNbO4 composite phase.

[0049] The preparation of the high-temperature-resistant aero-engine blade first formed the substrate. 70 g of TiAl alloy powder, 20 g of the prepared core-shell structure self-healing reinforcing phase, 10 g of SiC short fibers and 3 g of polycarbosilane solution were accurately weighed and mixed in a three-dimensional mixer at a speed of 150 r / min for 5 h to obtain a uniform mixture. The mixture was loaded into a BN release agent coated graphite mold, and a spark plasma sintering system was used to heat the mixture to 1180°C at a rate of 100°C / min under vacuum, while applying an axial pressure of 50 MPa and maintaining the temperature and pressure for 15 min to obtain a blade blank with a density of 99.2%. Then precise machining and coating preparation were performed. A five-axis numerical control machine tool was used to machine the precise blade geometry at a spindle speed of 8000 r / min and a feed speed of 800 mm / min. 15 g of rare earth niobate iron elastic protective compound powder, 5 g of HfO2-Y2O3 and 2 g of polymethyl methacrylate microspheres were uniformly mixed, and then a protective coating with a thickness of 200 μm was formed on the surface of the blade by an atmospheric plasma spraying device under the following process conditions: power 45 kW, spraying distance 120 mm, and powder feeding rate 25 g / min. The porosity of the coating was determined to be 12%. Finally, the coating blade was placed in a tube furnace and heated to 800°C at a rate of 10°C / min under argon protection, and then heat treated for 3 h and cooled in the furnace.

[0050] Example Two:

[0051] The difference between this embodiment and embodiment 1 is that, in the preparation of the core-shell structure self-healing reinforcing phase, 69 g of Ti powder and 33 g of B powder are ignited under argon protection to obtain a TiB2porous preform; the preform is placed in a CVD reactor to form a BN shell layer by reacting BCl3and NH3as precursors at 1200°C for 2.5 h. In the preparation of the rare earth niobate iron elastic protective compound, 70.5 g of Nd2O3, 54.5 g of Nb2O5, 12.8 g of Ta2O5and 13.0 g of ReO2are dissolved in dilute nitric acid, an excess of oxalic acid solution is added to generate a precipitate, which is then dispersed in deionized water after filtration and washing, a polyethylene glycol dispersant is added, and the precursor powder is obtained by spray drying; sintering at 1480°C in air atmosphere for 7 h. In the preparation of the high-temperature-resistant aero-engine blade, 68 g of TiAl alloy powder, 18 g of core-shell structure self-healing reinforcing phase, 12 g of SiC short fiber and 2.5 g of polycarbosilane solution are mixed, loaded into a graphite mold, and subjected to spark plasma sintering at 1200°C for 12 min to obtain a blade blank; after processing by a five-axis numerical control machine tool, 16 g of rare earth niobate iron elastic protective compound powder, 4 g of HfO2-Y2O3and 1.8 g of polymethyl methacrylate microspheres are mixed, and a protective coating with a thickness of 180 μm is formed by plasma spraying; heat treatment at 795°C under argon protection for 3.5 h.

[0052] Example Three

[0053] The difference between this embodiment and embodiment 1 is that, in the preparation of the core-shell structure self-healing reinforcing phase, 69 g of Ti powder and 33 g of B powder are ignited under argon protection to obtain a TiB2porous preform; the preform is placed in a CVD reactor to form a BN shell layer by reacting BCl3and NH3as precursors at 1200°C for 2.5 h. In the preparation of the rare earth niobate iron elastic protective compound, 70.5 g of Nd2O3, 54.5 g of Nb2O5, 12.8 g of Ta2O5and 13.0 g of ReO2are dissolved in dilute nitric acid, an excess of oxalic acid solution is added to generate a precipitate, which is then dispersed in deionized water after filtration and washing, a polyethylene glycol dispersant is added, and the precursor powder is obtained by spray drying; sintering at 1480°C in air atmosphere for 7 h. In the preparation of the high-temperature-resistant aero-engine blade, 68 g of TiAl alloy powder, 18 g of core-shell structure self-healing reinforcing phase, 12 g of SiC short fiber and 2.5 g of polycarbosilane solution are mixed, loaded into a graphite mold, and subjected to spark plasma sintering at 1200°C for 12 min to obtain a blade blank; after processing by a five-axis numerical control machine tool, 16 g of rare earth niobate iron elastic protective compound powder, 4 g of HfO2-Y2O3and 1.8 g of polymethyl methacrylate microspheres are mixed, and a protective coating with a thickness of 180 μm is formed by plasma spraying; heat treatment at 795°C under argon protection for 3.5 h.

[0054] Comparative Example 1

[0055] The difference between this comparative example and Example 1 is that 70 g of TiAl alloy powder, 20 g of ordinary TiB2 powder, 10 g of SiC short fibers and 3 g of polycarbosilane solution are mixed, loaded into a graphite mold, and subjected to spark plasma sintering at 1180°C for 15 min to obtain a vane blank; after processing by a five-axis numerical control machine tool, 15 g of ordinary yttrium-stabilized zirconia powder is mixed with 5 g of HfO2-Y2O3 and 2 g of polymethyl methacrylate microspheres, and a protective coating with a thickness of 200 μm is formed by plasma spraying; and heat treatment is performed at 800°C under argon protection for 3 h.

[0056] Comparative Example 2

[0057] The difference between this comparative example and Example 1 is that 70 g of TiAl alloy powder, 20 g of core-shell structure self-healing reinforcing phase prepared according to the method of Example 1, 10 g of SiC short fibers and 3 g of polycarbosilane solution are mixed, loaded into a graphite mold, and subjected to spark plasma sintering at 1180°C for 15 min to obtain a vane blank; after processing by a five-axis numerical control machine tool, 15 g of ordinary yttrium-stabilized zirconia powder is mixed with 5 g of HfO2-Y2O3 and 2 g of polymethyl methacrylate microspheres, and a protective coating with a thickness of 200 μm is formed by plasma spraying; and heat treatment is performed at 800°C under argon protection for 3 h.

[0058] Comparative Example 3

[0059] The difference between this comparative example and Example 1 is that 70 g of TiAl alloy powder, 20 g of ordinary TiB2 powder, 10 g of SiC short fibers and 3 g of polycarbosilane solution are mixed, loaded into a graphite mold, and subjected to spark plasma sintering at 1180°C for 15 min to obtain a vane blank; after processing by a five-axis numerical control machine tool, 15 g of rare earth niobate iron elastic protective compound prepared according to the method of Example 1 is mixed with 5 g of HfO2-Y2O3 and 2 g of polymethyl methacrylate microspheres, and a protective coating with a thickness of 200 μm is formed by plasma spraying; and heat treatment is performed at 800°C under argon protection for 3 h.

[0060] The properties of the high-temperature-resistant aero-engine blades obtained in Examples 1-3 and Comparative Examples 1-3 are tested according to the test specifications of national and industry standards.

[0061] High-temperature tensile testing is performed at 1100°C in an atmospheric environment using an electronic universal testing machine equipped with a high-temperature furnace. The sample is processed into a round bar with a diameter of 6 mm and a gauge length of 30 mm according to the standard, clamped, heated to the test temperature at a rate of 10°C / min, and kept for 20 min to ensure uniform temperature. Then, a tensile load is applied at a crosshead displacement rate of 1 mm / min until fracture, and the load-displacement curve is automatically recorded throughout the process, and the tensile strength, yield strength and elongation after fracture are calculated.

[0062] High temperature oxidation test was performed by placing the sample in a corundum crucible into a box furnace which was stabilized at 1100 °C, and continuously exposed for 500 h in static air environment. The sample was taken out every 100 h, cooled to room temperature, and the mass change was measured using an analytical balance with a precision of 0.1 mg. Finally, the mass change value per unit area was calculated as the oxidation resistance performance index.

[0063] Thermal shock test was performed by placing the sample in a 1100 °C box furnace for 10 min, then quickly transferred into deionized water at 25 °C for rapid cooling for 20 s. Subsequently, the sample was taken out, dried and the surface state was checked. This process was a cycle, and the operation was repeated until the coating was peeled off or visible cracks appeared on the substrate. The cycle number was recorded as the thermal shock life.

[0064] Thermal conductivity test was performed by laser flash method. The sample was processed into a circular sheet with a diameter of 12.7 mm and a thickness of 2.5 mm. The thermal diffusivity was measured in a vacuum environment at 1100 °C. The specific heat capacity was determined by differential scanning calorimetry. The material density was obtained by Archimedes drainage method. Finally, the thermal conductivity value was calculated.

[0065] High temperature stress rupture life test was performed by applying a constant tensile load of 100 MPa to the standard tensile sample in an air environment at 1100 °C. The total time from the start of loading to the fracture of the sample was recorded as the stress rupture life.

[0066] Fracture toughness test was performed by using compact tension sample. At room temperature, a sharp crack was pre-prepared by fatigue, and then a force was applied on a universal testing machine at a loading rate of 0.5 mm / min. The load-crack opening displacement curve was automatically recorded. The plane strain fracture toughness value was calculated according to the standard formula.

[0067] The performance test data are shown in Table 1.

[0068] Table 1 Performance test results

[0069]

[0070] The test results in Table 1 clearly show that the embodiments 1-3 of the present application solve the technical defects exposed by the comparative examples by using the synergistic system of the core-shell structure self-healing reinforcing phase and the rare earth niobate iron elastic protective compound. The comparative example 1 uses the traditional TiB2 reinforcing phase and the ordinary thermal barrier coating material, and performs poorly in the key indicators of high temperature tensile strength 310 MPa, oxidation weight gain 5.2 mg / cm2, etc., proving that the traditional material system has been unable to meet the harsh requirements of the new generation of aero-engine blades on high temperature resistance performance. Although the comparative example 2 introduces the core-shell structure self-healing reinforcing phase to improve the mechanical properties (high temperature tensile strength 335 MPa), the oxidation resistance performance (oxidation weight gain 4.8 mg / cm2) and the thermal protection performance (thermal conductivity 9.0 W / m·K) are limited, indicating that a single reinforcing phase is not enough to achieve a comprehensive performance breakthrough. The comparative example 3 uses a new protective coating but follows the traditional reinforcing phase, although the oxidation resistance performance is significantly improved (oxidation weight gain 1.5 mg / cm2), but the substrate strength (high temperature tensile strength 325 MPa) and the fracture toughness (11.5 MPa·m 1 / 2 ) are still significantly lower than the embodiments, confirming that the improvement of a single coating lacking effective reinforcing phase support has limitations. In contrast, the embodiment 1 achieves a comprehensive breakthrough in the key indicators of high temperature tensile strength 420 MPa, oxidation weight gain 0.8 mg / cm2, thermal conductivity 5.8 W / m·K, etc. by the synergistic effect of the two new compounds, the high temperature endurance life 508 hours is 3.3 times that of the comparative example 1, the thermal shock cycle number exceeds 100 times and there is no failure, and the fracture toughness 19.5 MPa·m 1 / 2 is significantly better than the comparative examples, fully verifying that the core-shell structure self-healing reinforcing phase effectively blocks crack propagation through the "crack pinning-healing" mechanism, and the rare earth niobate iron elastic protective compound realizes efficient thermal protection and stress regulation through the iron elastic domain switching and ultra-low thermal conductivity, and the multi-level protective system constructed by the synergistic effect of the two realizes the long-term stable service of the aero-engine blades in extreme environments.

Claims

1. A method for preparing a high-temperature resistant aero-engine blade, characterized in that the steps include... include: S1. TiAl alloy powder, core-shell self-healing reinforcing phase, SiC short fibers and polycarbosilane solution are mixed in a mixer to obtain a mixture; the mixture is loaded into a graphite mold and formed by spark plasma sintering, and held at a temperature of 1150-1250℃ to obtain a blade blank; S2. The blade blank is precision machined using a five-axis CNC machine tool to form the blade geometry; rare earth niobate iron elastic protective compound powder is mixed with HfO2-Y2O3 and polymethyl methacrylate microspheres, and sprayed on the blade surface using a plasma spraying equipment to form a blade with a protective coating. S3. Under argon protection at 790-810℃, the blades with protective coating are heat-treated and then cooled in the furnace. The preparation method of the core-shell structure self-healing reinforcement phase includes: A1, mixing Ti powder and B powder to obtain a mixture, molding the mixture into a green body; then igniting the green body under argon protection to obtain a TiB2 porous preform; A2, placing the TiB2 preform in a CVD reactor, using BCl3 and NH3 as precursors, and reacting at 1100-1300℃. The preparation method of the rare earth niobate iron elastic protective compound includes: B1, dissolving Nd2O3, Nb2O5, Ta2O5 and ReO2 in dilute nitric acid to form a mixed solution; adding oxalic acid solution to generate oxalate precipitate, filtering and washing with deionized water; redispersing the oxalate precipitate in deionized water, adding polyethylene glycol, and spray drying to obtain precursor powder; B2, sintering in air atmosphere at 1450-1550℃.

2. The method for preparing high-temperature resistant aero-engine blades according to claim 1, characterized in that, In step S2, the thickness of the protective coating on the blade is 150-300 μm.

3. The method for preparing high-temperature resistant aero-engine blades according to claim 1, characterized in that, In step S3, the heat treatment time is 2-4 hours.

4. The method for preparing high-temperature resistant aero-engine blades according to claim 1, characterized in that, In step A1, the mass ratio of Ti powder to B powder is 69:(32-34).

5. The method for preparing high-temperature resistant aero-engine blades according to claim 1, characterized in that, In step A2, the reaction time is 2-4 hours at 1100-1300℃.

6. The method for preparing high-temperature resistant aero-engine blades according to claim 1, characterized in that, In step B1, the mass ratio of Nd2O3, Nb2O5, Ta2O5 and ReO2 is (70.0-72.0):(53.5-55.0):(12.5-13.5):(12.0-13.5).

7. The method for preparing high-temperature resistant aero-engine blades according to claim 1, characterized in that, In step B2, the sintering time in an air atmosphere at 1450-1550℃ is 5-8 hours.

8. A high-temperature resistant aero-engine blade prepared by the method according to any one of claims 1-7, characterized in that, The raw materials include the following parts by weight: 60-75 parts TiAl alloy; 15-25 parts core-shell structure self-healing reinforcing phase; 10-20 parts rare earth niobate iron elastic protective compound; 5-15 parts SiC short fiber; 3-8 parts HfO2-Y2O3; 2-5 parts polycarbosilane; and 1-3 parts polymethyl methacrylate microspheres.

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