Gradient reinforced composite material gas turbine blade and laser three-dimensional forming preparation method

Abstract

The invention provides a laser three-dimensional forming preparation method for preparing a gradient reinforced composite material turbine blade of a gas turbine. The method comprises the following steps: ball-milling and mixing nickel-based alloy, nano silicon carbide and boron carbide powder according to a mass ratio of (92-96): (2-4): (2-4); a laser three-dimensional forming technology is adopted, a three-section gradient structure of a non-strengthening area (pure nickel base alloy), a transition area (gradient mixing) and a strengthening area (composite material) is constructed in the root-tip direction of the blade through a two-way powder feeding system, and near-net forming of the functional gradient strengthening turbine blade with the ceramic strengthening phase strengthening at the tip end of the blade and the root part maintaining toughness is achieved. The prepared turbine blade of the gas turbine has excellent high-temperature creep resistance, fatigue resistance, corrosion resistance and wear resistance, works for a long time in severe environments such as high temperature, high pressure, wear and corrosive media, has excellent high-temperature creep damage resistance, low-cycle fatigue damage resistance, corrosion damage resistance and friction damage resistance, and can effectively prolong the service life of the gas turbine.

Description

A gradient-reinforced composite gas turbine blade and a laser stereoforming method for its fabrication. Technical Field

[0001] This invention belongs to the field of laser stereolithography technology, specifically relating to a gradient-strengthened composite gas turbine blade and a laser stereolithography preparation method. Background Technology

[0002] Gas turbines are critical components in defense and energy power equipment, often referred to as the "crown jewel" of the equipment manufacturing industry. The hot-end components of heavy-duty gas turbines mainly include the combustion chamber, turbine blades, and turbine disk. Turbine blades, in particular, typically operate for extended periods in harsh environments with high temperatures, high pressures, abrasive surfaces, and corrosive media, making them susceptible to high-temperature creep damage, low-cycle fatigue damage, corrosion damage, and frictional damage, leading to failure. The operating temperature, load, and media characteristics of gas turbine blades have a complex impact on their damage; turbine blades operating under prolonged harsh conditions are prone to cracking, wear, and even fracture. Furthermore, compared to aero-turbine engine blades, gas turbine blades require significantly higher resistance to thermal corrosion. This is because fuel contains a large amount of sulfur, which readily generates SO2 and SO3 gases during combustion. Under high temperatures, these SO2 and SO3 react with elements such as Cr, Al, and Ti in the alloy to form brittle oxides, promoting crack initiation and growth, and resulting in a decline in material properties. As the performance requirements of gas turbines increase, the material performance requirements for their turbine blades also continuously rise, leading to increasingly higher production costs. However, compared with other developed countries, my country's gas turbine industry still has considerable room for development. It has not yet fully mastered the core design, manufacturing, maintenance, and control technologies of turbine blades, and exports of some advanced gas turbines to China remain restricted. Considering the extremely harsh service conditions of turbine blades, utilizing new materials and processes to manufacture gas turbine blades that meet performance requirements is particularly necessary for the development of my country's manufacturing industry. Summary of the Invention

[0003] The technical problem to be solved by this invention is to address the shortcomings of the prior art by providing a gradient-reinforced composite gas turbine blade and a laser stereolithography method for its preparation. This method utilizes laser stereolithography technology and a multi-channel powder delivery process to prepare the gradient-reinforced composite gas turbine blade. A high-temperature nickel-based alloy is used as the matrix, and the content of nano-silicon carbide ceramic particles and nano-boron carbide ceramic particles is distributed in a gradient according to three regions (non-reinforced region, transition region, and reinforced region) from the turbine blade root (the junction of the turbine blade and the bladed disk) to the turbine blade tip. This gas turbine blade exhibits excellent resistance to high-temperature creep, fatigue, corrosion, and wear. It can operate for extended periods in harsh environments such as high temperature, high pressure, abrasive, and corrosive media, demonstrating excellent resistance to high-temperature creep damage, low-cycle fatigue damage, corrosion damage, and friction damage, effectively extending the service life of the gas turbine.

[0004] To solve the above-mentioned technical problems, the technical solution adopted by the present invention is: a method for preparing gradient-reinforced composite gas turbine blades, characterized in that the method comprises: Step 1, drying nickel-based alloy powder, nano-silicon carbide ceramic powder, and nano-boron carbide ceramic powder respectively; Step 2, ball milling and mixing the dried nickel-based alloy powder, nano-silicon carbide ceramic powder, and nano-boron carbide ceramic powder in Step 1 at a mass ratio of (92-96):(2-4):(2-4) to obtain a mixed powder, and then drying the mixed powder; Step 3, loading the dried nickel-based alloy powder in Step 1 and the dried mixed powder in Step 2 into the multi-channel synchronous off-axis powder feeding system of a laser stereolithography equipment, using argon as a protective gas to preheat the substrate, and adjusting the first nozzle, the second nozzle, and the laser beam to be directly facing the substrate. After drying, nickel-based alloy powder is ejected from the first nozzle according to a preset powder feeding amount, and the dried mixed powder is ejected from the second nozzle according to a preset powder feeding amount. The laser beam follows the forming trajectory from the root of the turbine blade to the tip of the turbine blade, passes through the pre-established three-dimensional network model of the gas turbine blade, and is synchronously scanned by the equipment's built-in operating software according to the set program to form a light spot on the substrate, and the molten ejected nickel-based alloy powder and / or mixed powder are deposited layer by layer. The pre-established three-dimensional network model of the gas turbine blade is divided into three regions from the root of the turbine blade to the tip of the turbine blade, including a non-reinforced region, a transition region, and a reinforced region. The non-reinforced and reinforced regions are formed by nickel-based alloy powder deposition, the transition region is formed by nickel-based alloy powder and mixed powder deposition, and the reinforced region is formed by mixed powder deposition.

[0005] Preferably, the nickel-based alloy powder in step two is Inconel 718 nickel-based alloy powder with a particle size of 50 μm to 100 μm. Inconel 718 nickel-based alloy is prepared by mixing raw materials in the following weight percentages: Nb 4.92%, Al 0.54%, C 0.045%, Ti 0.97%, Cr 19.2%, Si 0.2%, Fe 18.1%, Mo 3.19%, Mn 0.04%, with the balance being Ni. The mass content of SiC in the nano-silicon carbide ceramic powder is >99%, and the particle size is 5 nm to 20 nm. The mass content of BC in the nano-boron carbide ceramic powder is >99%, and the particle size is 2 nm to 10 nm.

[0006] Preferably, the drying conditions in steps one and two are: drying in a high-temperature furnace at 150°C for 4 hours.

[0007] Preferably, the pre-established three-dimensional network model of the gas turbine blade is divided into n layers according to the laser beam scanning trajectory, where n is a multiple of 3. From the 1st layer to the n / 3rd layer, that is, from the root of the turbine blade to 2 / 3 of the distance from the tip of the turbine blade, is the non-reinforced region; from the n / 3+1st layer to the 2n / 3rd layer, that is, from 2 / 3 of the distance from the tip of the turbine blade to 1 / 3 of the distance from the tip of the turbine blade, is the transition region; and from the 2n / 3+1st layer to the nth layer, that is, from 1 / 3 of the distance from the tip of the blade to the tip of the turbine blade, is the reinforced region.

[0008] Preferably, when depositing to form the non-reinforced region, the powder feeding pressure is 0.6–0.8 MPa, the powder feeding voltage is 18–22 V, the scanning speed is 800–1000 mm / h, the laser power is 1100–1200 W, the powder feed rate of the first nozzle is 6–8 g / min, and the spot diameter is 2 mm; when depositing to form the transition region, the powder feeding pressure is 0.6–0.8 MPa, the powder feeding voltage is 18–22 V, and the scanning speed is 800–1000 mm / h. h, laser power is 1200-1300W, powder feed rate of the first nozzle is 3-4g / min, powder feed rate of the second nozzle is 3-4g / min, and spot diameter is 2mm; for deposition forming reinforcement region, powder feeding pressure is 0.6-0.8MPa, powder feeding voltage is 18-22V, scanning speed is 800-1000mm / h, laser power is 1300-1400W, powder feed rate of the second nozzle is 6-8g / min, and spot diameter is 2mm.

[0009] Preferably, the sum of the powder feed rate of the first nozzle in the non-reinforced region of deposition forming, the powder feed rate of the second nozzle in the reinforced region of deposition forming, and the powder feed rate of the first nozzle and the powder feed rate of the second nozzle in the transition region of deposition forming is equal.

[0010] Preferably, during the deposition of the transition region, the powder feed rate of the first nozzle and the powder feed rate of the second nozzle are the same.

[0011] Preferably, the pressure of the protective atmosphere argon gas in step three is 0.08–0.12 MPa.

[0012] Preferably, in step three, the substrate preheating temperature is 1400–1600°C and the preheating time is 0.5–1.5 h.

[0013] The present invention also discloses a gradient-strengthened composite gas turbine blade, characterized in that the gradient-strengthened composite gas turbine blade is prepared according to the method described in claims 1-9.

[0014] Compared with existing technologies, this invention has the following advantages: Utilizing laser stereolithography technology and a multi-channel powder delivery process, this invention prepares gradient-reinforced composite gas turbine blades. The composite material, with Inconel 718 nickel-based alloy as the matrix and nano-silicon carbide and nano-boron carbide ceramic particles as reinforcing phases, exhibits excellent resistance to high-temperature creep, fatigue, corrosion, and wear, fully meeting the service requirements of gas turbine blades. Furthermore, during the laser stereolithography process, the turbine blade root (the junction of the turbine blade and the blade disk) to the blade tip is divided into three regions based on the degree of damage during service. The content of the added nano-silicon carbide and nano-boron carbide ceramic particles as reinforcing phases is gradient-distributed, which can reduce the amount of nano-ceramics used and lower the manufacturing cost of the turbine blades.

[0015] The present invention discloses a method for preparing gradient-reinforced composite gas turbine blades, using nickel-based alloy Inconel 718 as the matrix material. This method effectively utilizes the corrosion resistance, oxidation resistance, good tensile and fatigue properties in high-temperature environments, resistance to stress corrosion under salt spray atmospheres, and excellent resistance to pitting corrosion, crevice corrosion, and intergranular corrosion in chloride media, inorganic acids, oxidizing and reducing environments.

[0016] Boron carbide ceramics are among the three hardest known materials (the other two being diamond and cubic boron nitride), characterized by low density, high strength, high-temperature stability, and good chemical stability. Silicon carbide ceramics possess high hardness, high corrosion resistance, and relatively high-temperature strength. Adding nano-silicon carbide and nano-boron carbide ceramics as reinforcing phases to the Inconel 718 nickel-based alloy matrix can further improve the high-temperature creep, fatigue, and corrosion resistance of gas turbine blades, significantly enhancing their high-temperature wear resistance. In particular, the pinning effect of nano-silicon carbide and nano-boron carbide ceramics can inhibit material cracking phenomena in nickel-based alloys, such as coarsening of the γ' phase structure, precipitation of brittle phases, creep porosity, and crack formation.

[0017] Turbine blades are located in the turbine section of a gas turbine. The hottest exhaust gas from the combustion chamber directly acts on the vicinity of the turbine blade tips. Due to engine start-up, shutdown, or sudden changes in state, high thermal stress is generated at the blade tips. Simultaneously, high-temperature creep elongation occurs, and wear occurs between the blade tips and the outer casing. Combined with the corrosion and oxidation effects of the high-temperature exhaust gas, premature cracking and other damage phenomena can occur at the turbine blade tips. Furthermore, the longer the service life, the more severe the wear damage at the blade tips, leading to the premature scrapping of many expensive turbine blades and causing significant economic losses. This invention creatively utilizes laser stereolithography technology to fabricate gradient-reinforced composite gas turbine blades. Based on physical studies of the damage caused by high-temperature creep, fatigue, corrosion, and wear during actual service, and the results of finite element analysis of the turbine blade service process, this invention addresses these issues. This invention creatively divides the turbine blade root (the junction of the turbine blade and the impeller) to the blade tip into three regions (non-reinforced region, transition region, and reinforced region) based on the degree of damage threat during service. The content of the added nano-ceramic reinforcing phase is distributed in a gradient along these three regions. Therefore, based on the damage situation of the gas turbine blade during service, this invention focuses on reinforcing the severely damaged blade tip region, performs general reinforcement on the less damaged middle section, and leaves the slightly damaged blade root region unreinforced. The gradually increasing content of the added nano-ceramic reinforcing phase along the regional gradient effectively reduces the impact of changes in thermal expansion coefficient, stiffness, and hardness caused by the addition of the nano-ceramic reinforcing phase on the overall performance of the turbine blade. Furthermore, because nano-ceramic reinforcing phases are expensive, the gradient distribution of their content reduces the amount used, thus lowering the manufacturing cost of the turbine blade.

[0018] The technical solution of the present invention will be further described in detail below with reference to the accompanying drawings and embodiments. Attached Figure Description

[0019] Figure 1 is a metallographic photograph of the gradient-reinforced composite gas turbine blade prepared in Example 1 of the present invention. Detailed Implementation

[0020] Example 1 This example discloses a method for preparing a gradient-reinforced composite gas turbine blade. The method is as follows: Step 1, drying nickel-based alloy powder, nano-silicon carbide ceramic powder, and nano-boron carbide ceramic powder in a high-temperature furnace at 150°C for 4 hours; Step 2, ball milling and mixing the dried nickel-based alloy powder, nano-silicon carbide ceramic powder, and nano-boron carbide ceramic powder in a mass ratio of 96:2:2 to obtain a mixed powder, and then drying the mixed powder in a high-temperature furnace at 150°C for 4 hours; The nickel-based alloy powder is Inconel 718 nickel-based alloy powder with a particle size of 50μm~100μm. Inconel 718 nickel-based alloy is prepared by mixing the following raw materials in weight percentage: Nb 4.92%, Al 0.54%, C 0.045%, Ti 0.97%, Cr 19.2%, Si 0.2%, Fe 18.1%, Mo 3.19%, Mn 0.04%, with the balance being Ni; the mass content of SiC in the nano-silicon carbide ceramic powder is >99%, and the particle size is 5nm~20nm; the mass content of BC in the nano-boron carbide ceramic powder is >99%, and the particle size is 2nm~10nm.

[0021] Step 3: Load the dried nickel-based alloy powder from Step 1 and the dried mixed powder from Step 2 into the multi-channel synchronous off-axis powder feeding system of the laser stereolithography equipment. Use argon as the protective gas, with a protective atmosphere pressure of 0.08 MPa, to preheat the substrate at 1400℃ for 1.5 hours. Adjust the first nozzle, the second nozzle, and the laser beam to face the substrate. The dried nickel-based alloy powder is ejected from the first nozzle according to the preset powder feeding rate, and the dried mixed powder is ejected from the second nozzle according to the preset powder feeding rate. The laser beam follows a forming trajectory from the root of the turbine blade to the tip of the turbine blade, passing through a pre-established three-dimensional network model of the gas turbine blade. The equipment's built-in operating software controls the synchronous scanning according to the set program to form a light spot on the substrate and cause the molten nickel to be ejected. The base alloy powder and / or mixed powder are deposited layer by layer to form the turbine blade. The pre-established three-dimensional network model of the gas turbine blade is divided into three regions from the root of the turbine blade to the tip of the turbine blade, which include a non-reinforced region, a transition region and a reinforced region. Specifically, the three-dimensional network model of the gas turbine blade is divided into n layers according to the laser beam scanning trajectory, where n is a multiple of 3. From the 1st layer to the n / 3rd layer, that is, from the root of the turbine blade (the junction of the turbine blade and the bladed disk) to 2 / 3 of the distance from the tip of the turbine blade, is the non-reinforced region; from the n / 3+1st layer to the 2n / 3rd layer, that is, from 2 / 3 of the distance from the tip of the turbine blade to 1 / 3 of the distance from the tip of the turbine blade, is the transition region; and from the 2n / 3+1st layer to the nth layer, that is, from 1 / 3 of the distance from the tip of the blade to the tip of the turbine blade, is the reinforced region.

[0022] In the non-reinforced and reinforced regions, nickel-based alloy powder is used for deposition during layer-by-layer deposition. The first nozzle is open, while the second nozzle remains closed. The powder feed pressure is 0.6 MPa, the powder feed voltage is 18 V, the scanning speed is 800 mm / h, the laser power is 1100 W, the powder feed rate of the first nozzle is 6 g / min, and the spot diameter is 2 mm. In the transition region, nickel-based alloy powder and mixed powder are used for co-deposition during deposition. The second nozzle is open, while the first nozzle remains open. The powder feed pressure is 0.6 MPa. The powder feeding voltage is 18V, the scanning speed is 800mm / h, the laser power is 1200W, the powder feeding rate of the first nozzle is 3g / min, the powder feeding rate of the second nozzle is 3g / min, and the spot diameter is 2mm. During the layer-by-layer deposition and forming process of the enhanced region, mixed powder deposition and forming is adopted. The first nozzle is closed, while the second nozzle is kept open. The powder feeding pressure is 0.6MPa, the powder feeding voltage is 18V, the scanning speed is 800mm / h, the laser power is 1300W, the powder feeding rate of the second nozzle is 6g / min, and the spot diameter is 2mm.

[0023] The automatic laser stereolithography equipment is the BLT-400 laser stereolithography equipment, manufactured by Xi'an Bright Laser Technologies Co., Ltd. Synchronous powder feeding laser stereolithography mainly uses a laser beam to synchronously irradiate the powder, causing it to melt, deposit, and solidify, achieving layer-by-layer stacking to prepare the desired material.

[0024] Figure 1 is a metallographic photograph of the gradient-strengthened composite gas turbine blade prepared in Example 1. It can be clearly observed from the figure that the microstructure of the turbine blade is uniform and there are no structural defects, indicating that the gradient-strengthened composite gas turbine blade prepared by laser stereolithography in this example has a good microstructure and excellent mechanical properties.

[0025] Example 2 This example discloses a method for preparing a gradient-reinforced composite gas turbine blade. The method is as follows: Step 1, drying nickel-based alloy powder, nano-silicon carbide ceramic powder, and nano-boron carbide ceramic powder in a high-temperature furnace at 150°C for 4 hours; Step 2, ball milling and mixing the dried nickel-based alloy powder, nano-silicon carbide ceramic powder, and nano-boron carbide ceramic powder in a mass ratio of 94:3:3 to obtain a mixed powder, and then drying the mixed powder in a high-temperature furnace at 150°C for 4 hours; the nickel-based alloy powder is Inconel 718 nickel-based alloy powder with a particle size of 50μm~100μm, and Inconel 718 nickel-based alloy is prepared by mixing the following raw materials in weight percentage: Nb 4.92%, Al 0.54%, C 0.045%, Ti 0.97%, Cr 19.2%, Si 0.2%, Fe 18.1%, Mo 3.19%, Mn 0.04%, with the balance being Ni; the mass content of SiC in the nano-silicon carbide ceramic powder is >99%, and the particle size is 5nm~20nm; the mass content of BC in the nano-boron carbide ceramic powder is >99%, and the particle size is 2nm~10nm.

[0026] Step 3: The dried nickel-based alloy powder from Step 1 and the dried mixed powder from Step 2 are loaded into the multi-channel synchronous off-axis powder feeding system of the laser stereolithography equipment. Argon is used as the protective gas, with a pressure of 0.1 MPa. The substrate is preheated at 1500℃ for 1 hour. The first nozzle, the second nozzle, and the laser beam are adjusted to face the substrate. The dried nickel-based alloy powder is ejected from the first nozzle according to the preset powder feeding rate, and the dried mixed powder is ejected from the second nozzle according to the preset powder feeding rate. The laser beam follows a forming trajectory from the root of the turbine blade to the tip of the turbine blade. Through a pre-established three-dimensional network model of the gas turbine blade, the equipment's built-in operating software controls the synchronous scanning according to the set program to form a light spot on the substrate. The molten nickel-based alloy powder and / or mixed powder are deposited layer by layer to form the substrate. The pre-established three-dimensional network model of the gas turbine blade extends from the root of the turbine blade to the tip of the turbine blade. The turbine blade tip is divided into three regions: a non-reinforced region, a transition region, and a reinforced region. Specifically, the three-dimensional network model of the gas turbine blade is divided into n layers according to the laser beam scanning trajectory, where n is a multiple of 3. From layer 1 to layer n / 3, that is, from the root of the turbine blade (where the turbine blade meets the bladed disk) to 2 / 3 of the distance from the turbine blade tip, is the non-reinforced region; from layer n / 3+1 to layer 2n / 3, that is, from 2 / 3 of the distance from the turbine blade tip to 1 / 3 of the distance from the turbine blade tip, is the transition region; and from layer 2n / 3+1 to layer n, that is, from 1 / 3 of the distance from the blade tip to the turbine blade tip, is the reinforced region.

[0027] In the non-reinforced and reinforced regions, layer-by-layer deposition is performed using nickel-based alloy powder. The first nozzle is open, while the second nozzle remains closed. The powder feed pressure is 0.7 MPa, the powder feed voltage is 20 V, the scanning speed is 900 mm / h, the laser power is 1150 W, the powder feed rate from the first nozzle is 7 g / min, and the spot diameter is 2 mm. In the transition regions, layer-by-layer deposition is performed using a combination of nickel-based alloy powder and mixed powder. The second nozzle is open, while the first nozzle remains open. The powder feed pressure is 0.7 MPa, the powder feed voltage is 20 V, the scanning speed is 900 mm / h, the laser power is 1250 W, the powder feed rate from the first nozzle is 3.5 g / min, and the powder feed rate from the second nozzle is 3.5 g / min. The laser beam speed is 7 g / min, the spot diameter is 2 mm; during the layer-by-layer deposition and forming process of the enhanced region, mixed powder deposition and forming is adopted, the first nozzle is closed, the second nozzle is kept open, the powder feeding pressure is 0.7 MPa, the powder feeding voltage is 20 V, the scanning speed is 900 mm / h, the laser power is 1350 W, the powder feeding amount of the second nozzle is 7 g / min, and the spot diameter is 2 mm.

[0028] Example 3 This example discloses a method for preparing a gradient-reinforced composite gas turbine blade. The method is as follows: Step 1, drying nickel-based alloy powder, nano-silicon carbide ceramic powder, and nano-boron carbide ceramic powder in a high-temperature furnace at 150°C for 4 hours; Step 2, ball milling and mixing the dried nickel-based alloy powder, nano-silicon carbide ceramic powder, and nano-boron carbide ceramic powder in a mass ratio of (92-96):(2-4):(2-4) to obtain a mixed powder, and then drying the mixed powder in a high-temperature furnace at 150°C for 4 hours; The nickel-based alloy powder is nickel-based alloy Inconel 718 powder with a particle size of 50μm~100μm. The nickel-based alloy Inconel 718 is prepared by mixing the following raw materials in weight percentage: Nb 4.92%, Al 0.54%, C 0.045%, Ti 0.97%, Cr 19.2%, Si 0.2%, Fe 18.1%, Mo 3.19%, Mn 0.04%, with the balance being Ni; the mass content of SiC in the nano-silicon carbide ceramic powder is >99%, and the particle size is 5nm~20nm; the mass content of BC in the nano-boron carbide ceramic powder is >99%, and the particle size is 2nm~10nm.

[0029] Step 3: The dried nickel-based alloy powder from Step 1 and the dried mixed powder from Step 2 are loaded into the multi-channel synchronous off-axis powder feeding system of the laser stereolithography equipment. Argon is used as the protective gas, with a pressure of 0.12 MPa. The substrate is preheated at 1600℃ for 0.5 hours. The first nozzle, the second nozzle, and the laser beam are adjusted to face the substrate. The dried nickel-based alloy powder is ejected from the first nozzle according to the preset powder feeding rate, and the dried mixed powder is ejected from the second nozzle according to the preset powder feeding rate. The laser beam follows a forming trajectory from the root of the turbine blade to the tip of the turbine blade, passing through a pre-established three-dimensional network model of the gas turbine blade. The equipment's built-in operating software controls the synchronous scanning according to the set program to form a light spot on the substrate, and the molten nickel-based alloy powder and / or mixed powder are deposited layer by layer. The pre-established three-dimensional network model of the gas turbine blade extends from the root of the turbine blade to the tip of the turbine blade. The turbine blade tip is divided into three regions: a non-reinforced region, a transition region, and a reinforced region. Specifically, the three-dimensional network model of the gas turbine blade is divided into n layers according to the laser beam scanning trajectory, where n is a multiple of 3. From layer 1 to layer n / 3, that is, from the root of the turbine blade (where the turbine blade meets the bladed disk) to 2 / 3 of the distance from the turbine blade tip, is the non-reinforced region; from layer n / 3+1 to layer 2n / 3, that is, from 2 / 3 of the distance from the turbine blade tip to 1 / 3 of the distance from the turbine blade tip, is the transition region; and from layer 2n / 3+1 to layer n, that is, from 1 / 3 of the distance from the blade tip to the turbine blade tip, is the reinforced region.

[0030] In the non-reinforced and reinforced regions, nickel-based alloy powder is used for deposition. The first nozzle is open, while the second nozzle remains closed. The powder feed pressure is 0.8 MPa, the powder feed voltage is 22 V, the scanning speed is 1000 mm / h, the laser power is 1200 W, the powder feed rate of the first nozzle is 8 g / min, and the spot diameter is 2 mm. In the transition region, nickel-based alloy powder and mixed powder are used for deposition. The second nozzle is open, while the first nozzle remains open. The powder feed pressure is 0.8 MPa, the powder feed voltage is 22 V, and the scanning speed is 1000 mm / h. The laser power is 1200 W, the powder feed rate of the first nozzle is 8 g / min, and the spot diameter is 2 mm. The scanning speed is 1000 mm / h, the laser power is 1300 W, the powder feed rate of the first nozzle is 4 g / min, the powder feed rate of the second nozzle is 4 g / min, and the spot diameter is 2 mm. During the layer-by-layer deposition and forming process of the enhanced region, mixed powder deposition and forming is adopted. The first nozzle is closed, the second nozzle is kept open, the powder feed pressure is 0.8 MPa, the powder feed voltage is 22 V, the scanning speed is 1000 mm / h, the laser power is 1400 W, the powder feed rate of the second nozzle is 8 g / min, and the spot diameter is 2 mm.

[0031] The average corrosion rate of the corrosion samples from the reinforced region of the turbine blade tip in Examples 1-3 was determined using weight loss analysis. The analytical method was based on the general principles of corrosion testing of metals and alloys (GB / T 19291-2003). The results are shown in Table 1. Table 1 shows the average corrosion rate of the corrosion samples from the reinforced region of the turbine blade tip after immersion in a simulated corrosion solution for 30, 60, and 90 days, respectively. As can be observed from Table 1, under the same corrosion conditions, the average corrosion rate of the corrosion samples from the reinforced region of the turbine blade tip prepared by this invention is significantly lower than that of the corrosion samples from the unreinforced region of the turbine blade. Therefore, the gradient-reinforced composite gas turbine blade prepared by this invention can effectively improve corrosion resistance.

[0032] Table 1. Corrosion resistance of the reinforced region at the tip of the gradient-reinforced composite gas turbine blade prepared according to the present invention. Friction and wear tests were conducted on the wear samples of the reinforced region at the tip of the turbine blades in Examples 1-3 using a friction and wear testing machine. A load of 20 N was applied at a rotational speed of 1200 r / min. Table 2 shows the wear amount of the wear samples in the reinforced region of the turbine blade tip after 20 h, 40 h, and 60 h of wear, respectively. From the test data in Table 2, it can be observed that under the same friction and wear test conditions, the wear amount of the prepared wear samples in the reinforced region of the turbine blade tip is significantly lower than that of the wear samples in the unreinforced region of the turbine blade. Therefore, the gradient-reinforced composite gas turbine blade prepared in this invention can effectively improve wear resistance.

[0033] Table 2. Wear resistance of the reinforced region at the tip of the gradient-reinforced composite material for gas turbine blades prepared according to the present invention. The above description is merely a preferred embodiment of the present invention and is not intended to limit the invention in any way. Any simple modifications, alterations, and equivalent changes made to the above embodiments based on the inventive essence shall still fall within the protection scope of the present invention.

Claims

1. A method for preparing gradient-reinforced composite gas turbine blades, characterized in that, The method is as follows: Step 1: Dry the nickel-based alloy powder, nano-silicon carbide ceramic powder, and nano-boron carbide ceramic powder separately; Step 2: Ball mill and mix the dried nickel-based alloy powder, nano-silicon carbide ceramic powder, and nano-boron carbide ceramic powder in a mass ratio of (92-96):(2-4):(2-4) to obtain a mixed powder, and then dry the mixed powder; Step 3: Load the dried nickel-based alloy powder in Step 1 and the dried mixed powder in Step 2 into the multi-channel synchronous off-axis powder feeding system of the laser stereolithography equipment, use argon as a protective gas to preheat the substrate, adjust the first nozzle, the second nozzle, and the laser beam to face the substrate, and after drying, the nickel-based alloy powder is ejected from the first nozzle according to the preset powder feeding amount, and then dried. After drying, the mixed powder is ejected through the second nozzle according to the preset powder feeding amount. The laser beam follows the forming trajectory from the root of the turbine blade to the tip of the turbine blade. It passes through the pre-established three-dimensional network model of the gas turbine blade and is synchronously scanned by the equipment's built-in operating software according to the set program to form a light spot on the substrate. The molten nickel-based alloy powder and / or mixed powder are deposited layer by layer to form the shape. The pre-established three-dimensional network model of the gas turbine blade is divided into three regions from the root of the turbine blade to the tip of the turbine blade, including a non-reinforced region, a transition region, and a reinforced region. The non-reinforced and reinforced regions are formed by depositing nickel-based alloy powder, the transition region is formed by depositing nickel-based alloy powder and mixed powder together, and the reinforced region is formed by depositing mixed powder.

2. The method according to claim 1, characterized in that, The nickel-based alloy powder mentioned in step two is Inconel 718 nickel-based alloy powder with a particle size of 50 μm to 100 μm. Inconel 718 nickel-based alloy is made by mixing raw materials in the following weight percentages: Nb 4.92%, Al 0.54%, C 0.045%, Ti 0.97%, Cr 19.2%, Si 0.2%, Fe 18.1%, Mo 3.19%, Mn 0.04%, with the balance being Ni. The mass content of SiC in the nano-silicon carbide ceramic powder is >99%, and the particle size is 5 nm to 20 nm. The mass content of BC in the nano-boron carbide ceramic powder is >99%, and the particle size is 2 nm to 10 nm.

3. The method according to claim 1, characterized in that, The drying conditions in steps one and two are: drying in a high-temperature furnace at 150℃ for 4 hours.

4. The method according to claim 1, characterized in that, The pre-established three-dimensional network model of the gas turbine blade is divided into n layers according to the laser beam scanning trajectory, where n is a multiple of 3. From the 1st layer to the n / 3rd layer, that is, from the root of the turbine blade to 2 / 3 of the distance from the tip of the turbine blade, is the non-reinforced region; from the n / 3+1st layer to the 2n / 3rd layer, that is, from 2 / 3 of the distance from the tip of the turbine blade to 1 / 3 of the distance from the tip of the turbine blade, is the transition region; and from the 2n / 3+1st layer to the nth layer, that is, from 1 / 3 of the distance from the tip of the blade to the tip of the turbine blade, is the reinforced region.

5. The method according to claim 1, characterized in that, When depositing to form the non-reinforced region, the powder feeding pressure is 0.6–0.8 MPa, the powder feeding voltage is 18–22 V, the scanning speed is 800–1000 mm / h, the laser power is 1100–1200 W, the powder feeding rate of the first nozzle is 6–8 g / min, and the spot diameter is 2 mm; when depositing to form the transition region, the powder feeding pressure is 0.6–0.8 MPa, the powder feeding voltage is 18–22 V, the scanning speed is 800–1000 mm / h, the laser power is 1200–1300 W, the powder feeding rate of the first nozzle is 3–4 g / min, the powder feeding rate of the second nozzle is 3–4 g / min, and the spot diameter is 2 mm. The deposition formation reinforcement area has a powder feeding pressure of 0.6–0.8 MPa, a powder feeding voltage of 18–22 V, a scanning speed of 800–1000 mm / h, a laser power of 1300–1400 W, a second nozzle powder feeding rate of 6–8 g / min, and a spot diameter of 2 mm.

6. The method according to claim 1, characterized in that, The sum of the powder feed rate of the first nozzle in the non-reinforced region of deposition forming, the powder feed rate of the second nozzle in the reinforced region of deposition forming, and the powder feed rate of the first nozzle and the powder feed rate of the second nozzle in the transition region of deposition forming is equal.

7. The method according to claim 6, characterized in that, During the deposition and forming transition zone, the powder feed rate of the first nozzle and the powder feed rate of the second nozzle are the same.

8. The method according to claim 1, characterized in that, In step three, the pressure of the protective atmosphere argon gas is 0.08–0.12 MPa.

9. The method according to claim 1, characterized in that, In step three, the substrate preheating temperature is 1400–1600℃, and the preheating time is 0.5–1.5 h.

10. A gradient-reinforced composite gas turbine blade, characterized in that, Gradient-reinforced composite gas turbine blades prepared according to the method described in claims 1-9.

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