Turbine blade assembly, airfoil module therefor, and method of manufacture

Abstract

The application discloses a turbine blade assembly and an airfoil module thereof. The airfoil module comprises an airfoil body extending in a spanwise direction of a turbine engine, comprising an outer curved surface formed on an outer side thereof and a cooling channel network arranged on an inner side thereof; and a flange located at an end of the airfoil body and extending transversely relative to the spanwise direction, so as to form a first joint surface on a surface of the flange. The airfoil module can be integrally formed by using an additive manufacturing process, and when the airfoil module is assembled with a platform module to form the turbine blade assembly, the flange of the airfoil module enlarges the area of the first joint surface, so that the joint strength and the uniformity of thermal stress distribution are improved, and the risk of interface peeling under high-temperature working conditions is effectively alleviated.

Description

Technical Field

[0001] This invention relates to the field of gas turbine engine manufacturing technology, and more specifically, to turbine blade assemblies, their airfoil modules, and manufacturing methods. Background Technology

[0002] Gas turbine engines typically consist of multiple turbine stages arranged sequentially, each stage generally containing stator blades and rotor blades. The stator blades guide the flow of high-temperature gas and regulate its flow angle and velocity distribution, while the rotor blades generate torque and output power under the influence of the gas. During operation, these blades are subjected to high-temperature gas erosion, centrifugal loads, and cyclic thermo-mechanical coupling stresses, and their service environment is characterized by high temperature, high speed, and significant temperature gradients. Therefore, the blade structure not only needs to possess high-temperature strength and creep resistance but also needs to maintain dimensional stability and cooling efficiency throughout its entire life cycle.

[0003] To meet the demands of high-temperature operation, turbine blades typically incorporate complex internal cooling networks. In the high-pressure turbine region, the internal structure may include multi-stage cooling structures such as serpentine channels, impingement cooling chambers, steering chambers, throttling features, and film cooling holes to reduce the temperature of the metal matrix and regulate the thermal gradient distribution. The geometry, cross-sectional dimensions, and surface condition of the internal cooling channels directly affect the flow pressure drop, heat transfer efficiency, and local thermal stress levels. Although low-pressure turbine blades operate at relatively lower temperatures, they may still include internal cavities and reinforcing structures to maintain structural stiffness and aerodynamic stability over a wide range of operating conditions.

[0004] In existing technologies, turbine blades can be manufactured using various methods, including integral casting, machining, additive manufacturing, and assembly after modular manufacturing. In high-temperature critical load-bearing regions, integral forming processes have a wide range of applications. For example, nickel-based superalloy blades can be formed through precision casting, and directional solidification or single-crystal processes can be combined to improve high-temperature microstructure stability. This type of integral structure is beneficial for maintaining material continuity and the integrity of the load transfer path. However, integral casting typically relies on ceramic cores to form internal cooling channels, and its geometric complexity and dimensional accuracy are limited by core manufacturing capabilities.

[0005] Additive manufacturing offers high geometrical design freedom for internal cooling structures, enabling the formation of complex three-dimensional channel networks without the need for ceramic cores. However, the layer-by-layer forming characteristic can lead to fluctuations in the surface roughness, cross-sectional consistency, and microstructure of internal channels. For integrally manufactured additive turbine blade components, the physical accessibility of some high-curvature or narrow internal regions is limited, increasing the difficulty of post-processing and quality inspection of internal surfaces. These factors may affect internal flow properties and local stress distribution, thus impacting the long-term service reliability of the blade under high-temperature thermo-mechanical cyclic loading conditions. Furthermore, the flexibility for subsequent maintenance or repair is relatively limited when defects or performance degradation occur in the internal regions of an integral structure.

[0006] Furthermore, existing technologies also include methods that divide the blades into multiple modules, manufacture them separately, and then form an integral structure through welding, brazing, or diffusion bonding. This type of modular structure improves manufacturing flexibility to some extent and allows for differentiated manufacturing paths in different areas. However, the connection interface must simultaneously meet requirements for structural load-bearing capacity, sealing performance, and dimensional accuracy under high-temperature thermo-mechanical cyclic loading. Factors such as interface gap control, thermal expansion matching, and the flow behavior of the connecting materials can all affect long-term reliability, making the connection area a critical constraint on the overall structure.

[0007] The above problems essentially reflect that in the design and manufacturing process of turbine blades, there are mutual constraints among structural architecture, manufacturing path selection, and the combination of different components, and there is still a lack of an overall technical solution that can achieve synergistic optimization among various performance and design requirements. Summary of the Invention

[0008] In view of the above problems, the purpose of the present invention is to provide a turbine blade assembly and its airfoil module, wherein the airfoil module and the platform module are manufactured separately and assembled together by brazing process, and the flange of the airfoil module is used to optimize the stress and interface state of the connection area, thereby improving the strength stability and durability of the turbine blade under high temperature thermo-mechanical cyclic load.

[0009] According to one aspect of the invention, an airfoil module for a turbine engine is provided, comprising: an airfoil body extending along the spanwise direction of the turbine engine, including an outer curved surface formed on its outer side and a cooling channel network disposed therein; and a flange located at an end of the airfoil body, extending laterally relative to the spanwise direction, thereby forming a first mating surface on the surface of the flange.

[0010] Optionally, the outer curved surface of the airfoil body is formed with a plurality of cooling holes that communicate with the cooling channel network.

[0011] Optionally, the surface of the flange is formed with a first microgroove structure for receiving solder.

[0012] Optionally, the depth of the first microgroove structure is 10 micrometers to 200 micrometers.

[0013] Optionally, the position of the first microgroove structure is offset from the steep transition zone of the main load-bearing stress peak region and / or thermal gradient region of the airfoil module.

[0014] Optionally, the first microgroove structure is located away from the blade root of the airfoil module.

[0015] Optionally, the first microgroove structure includes a plurality of microgrooves arranged in an array.

[0016] Optionally, the first microgroove structure includes a plurality of nested annular grooves.

[0017] Optionally, the first microgroove structure includes multiple nested arc-shaped grooves.

[0018] Optionally, an outer blocking structure is formed on the surface of the flange around the first microgroove structure to limit radial overflow of the solder during the high-temperature melting process.

[0019] Optionally, the outer blocking structure is an annular groove.

[0020] Optionally, the cooling channel network has inlets and / or outlets formed on the surface of the flange.

[0021] Optionally, an inner blocking structure is formed on the surface of the flange around the inlet and / or outlet to prevent the brazing filler metal from covering the inlet and / or outlet during the high-temperature melting process.

[0022] Optionally, the inner blocking structure is an annular groove.

[0023] Optionally, the annular groove includes an annular groove and a flow-blocking material that partially fills the annular groove.

[0024] Optionally, the airfoil body and the flange are integrally formed using an additive manufacturing process.

[0025] Optionally, the flange has a contour shape that is conformal or non-conformal to the end of the airfoil body.

[0026] Optionally, the flange sidewall surface is formed with a second microgroove structure for receiving brazing filler metal.

[0027] According to another aspect of the present invention, a turbine blade assembly is provided, comprising: at least one airfoil module as described above; and a platform module having a second mating surface that matches a first mating surface of the at least one airfoil module.

[0028] Optionally, the second mating surface of the platform module is a flat surface or a recessed surface of the platform module.

[0029] Optionally, the flange of the airfoil module is embedded in the recessed surface of the platform module.

[0030] Optionally, the platform module includes a mechanical positioning structure for precise alignment with an external wheel.

[0031] Optionally, the mechanical positioning structure includes at least one of dovetail tenon, tongue tenon, groove tenon, and keyway.

[0032] Optionally, the turbine blade assembly is a rotor blade assembly, and the respective flanges of the at least one airfoil module are located at the bottom end of the airfoil body.

[0033] Optionally, the turbine blade assembly is a stator blade assembly, and the at least one airfoil module includes a first flange located at the top of the airfoil body and a second flange located at the bottom of the airfoil body.

[0034] Optionally, the stator blade assembly is a sector segment, and the at least one airfoil module includes multiple airfoil modules that share the platform module, the multiple airfoil modules being distributed circumferentially along the sector segment.

[0035] Optionally, the platform module is manufactured using traditional manufacturing processes.

[0036] According to another aspect of the present invention, a method for manufacturing a turbine blade assembly is provided, comprising: forming at least one airfoil module using an additive manufacturing process; forming a platform module using a conventional manufacturing process; and joining the at least one airfoil module to the platform module using a brazing process, wherein the at least one airfoil module comprises: an airfoil body extending along the spanwise direction of the turbine engine, including an outer curved surface formed on its outer side and a cooling channel network disposed therein; and a flange located at the end of the airfoil body, extending laterally relative to the spanwise direction, thereby forming a first mating surface on the surface of the flange, the platform module being provided with a second mating surface matching the first mating surface of the at least one airfoil module, the flange of the at least one airfoil module increasing the area of ​​the first mating surface to improve the bonding strength and thermal stress distribution uniformity.

[0037] Optionally, the surface of the flange is formed with a first microgroove structure for receiving brazing filler metal. During the brazing process, the first microgroove structure guides the brazing filler metal to be evenly distributed along the interface between the at least one airfoil module and the platform module.

[0038] Optionally, the sidewall surface of the flange is formed with a second microgroove structure for accommodating the brazing filler metal. During the brazing process, the first microgroove structure guides the brazing filler metal to be evenly distributed along the interface between the at least one airfoil module and the platform module.

[0039] Optionally, an outer blocking structure is formed on the surface of the flange surrounding the first microgroove structure, and during the brazing process, the outer blocking structure restricts the radial overflow of the brazing filler metal during the high-temperature melting process.

[0040] Optionally, the cooling channel network has an inlet and / or outlet formed on the surface of the flange, and an inner blocking structure is formed on the surface of the flange surrounding the inlet and / or outlet. During the brazing process, the inner blocking structure prevents the brazing filler metal from covering the inlet and / or outlet during high-temperature melting.

[0041] Optionally, the additive manufacturing process includes at least one of selective laser melting, electron beam melting, directional energy deposition, or fused deposition modeling.

[0042] Optionally, the conventional manufacturing process includes at least one forming process selected from casting, forging, powder metallurgy, or precision casting, as well as a machining process.

[0043] Optionally, prior to the brazing process, the process further includes hot isostatic pressing of the at least one airfoil module.

[0044] Optionally, both the at least one airfoil module and the platform module are made of nickel-based high-temperature alloys, and the brazing process uses a nickel-based low-melting-point brazing filler metal containing nickel, chromium, boron and silicon.

[0045] Compared with the prior art, the turbine blade assembly and its airfoil module provided by the present invention are manufactured separately from the platform module and assembled together by brazing.

[0046] This invention utilizes a modular design, enabling the independent manufacture of airfoil and platform modules. This allows for the selection of optimal manufacturing processes and material systems to meet different functional requirements. For example, the airfoil module can be integrally formed using additive manufacturing processes to allow for the free design of its complex internal cooling channel network; while the platform module can be manufactured using traditional processes such as casting and forging to optimize its structural strength, dimensional stability, or oxidation resistance. This separate manufacturing approach not only improves the performance of each module but also significantly reduces the overall risk of scrapping. When a module (such as a high-value platform module) develops a localized defect, it is not necessary to scrap the entire blade assembly; only the module needs to be replaced or repaired, thereby achieving cost savings and efficient resource utilization.

[0047] In a preferred embodiment, the present invention provides a laterally extending flange at the end of the airfoil module, significantly increasing the bonding area with the platform module. This design provides a larger metallurgical bonding area during brazing, which not only directly improves the static strength and fatigue resistance of the joint, but also facilitates a more uniform distribution of interfacial thermal stress under high-temperature thermal cycling loads. This effectively mitigates the risk of interfacial delamination caused by differences in thermal expansion coefficients or temperature gradients, thereby ensuring the long-term stability of the connection interface under harsh service environments.

[0048] In a preferred embodiment, a microgroove structure is formed on the mating surface of the flange, creating spatial concave-convex geometric features in the interface region, which can serve as a channel for receiving and guiding the brazing filler metal. During brazing, these microgroove structures can guide the brazing filler metal to spread and fill more evenly along the mating interface, forming a continuous metallurgical bonding layer in multiple directions. This helps to form a dense, continuous, and uniformly thick brazed seam, reducing defects such as incomplete penetration and porosity, and further ensuring the connection quality. In particular, when the microgroove structure is designed in the form of an array, annular groove, or arc-shaped groove, the storage and capillary flow path of the brazing filler metal can be optimized.

[0049] In a preferred embodiment, the microgroove structure is positioned away from the peak region of the airfoil module's main load-bearing stress and / or the abrupt change region of the thermal gradient. For example, the microgroove structure is located away from the high-stress concentration region at the blade root, while also avoiding the high-temperature and high-temperature gradient regions on the aerodynamic outer surface. This prevents the brazing interface from overlapping with the peak of the main load-bearing stress (and / / the abrupt change region of the thermal gradient), thereby helping to improve the stability and durability of the interface under high-temperature thermo-mechanical cyclic loading.

[0050] In a preferred embodiment, outer and inner blocking structures are positioned near the microgroove structure to provide precise control over the solder flow. The outer blocking structure limits the overflow of molten solder into undesirable areas around the joint surface at high temperatures, ensuring a neat solder joint. The inner blocking structure, particularly those positioned around the inlet or outlet of the cooling channel, effectively prevents solder from clogging these critical flow characteristics, thus ensuring unobstructed internal cooling network and reliable cooling performance while achieving a robust connection. The blocking structure can be implemented using annular grooves partially filled with flow-blocking material, enhancing the flow-blocking effect.

[0051] In a preferred embodiment, the stator blade assembly can employ a sector-shaped segment design, where a shared platform module is connected to multiple airfoil modules. This structure reduces the number of sealing seams between platforms in the overall engine assembly, improving the aerodynamic continuity and sealing of the flow channels. Simultaneously, the modular sector-shaped segments facilitate partial replacement during maintenance, further enhancing the maintainability and lifecycle cost-effectiveness of the assembly.

[0052] The manufacturing method provided by this invention systematically integrates additive manufacturing, traditional forming processes, hot isostatic pressing (HIP), and controlled brazing. This method fully leverages the flexibility of additive manufacturing in forming complex geometries, ensures the densification and performance optimization of additive-manufactured components through HIP, and achieves high-strength, high-sealing connections between modules through precision brazing. This hybrid manufacturing approach provides an efficient and repeatable solution for producing high-performance, high-reliability turbine blade assemblies.

[0053] In summary, the turbine blade assembly and its airfoil module of the present invention, through innovative flange structure design combined with modular manufacturing concept, as well as optional microgrooves and blocking features, have achieved significant technical progress in improving connection strength, optimizing thermal stress management, ensuring the integrity of cooling function, and improving manufacturing economy and maintainability. It is particularly suitable for high-pressure turbine and low-pressure turbine blades of gas turbine engines with extremely high requirements for reliability and life. Attached Figure Description

[0054] Figure 1 A schematic cross-sectional view showing the basic structure of a turbine engine;

[0055] Figure 2 A perspective view of a rotor blade assembly according to a first embodiment of the present invention is shown;

[0056] Figure 3 Show Figure 2 The diagram shows the separated state of the airfoil module and platform module of the rotor blade assembly.

[0057] Figure 4A perspective view showing the basic structure of the stator blade assembly according to a second embodiment of the present invention;

[0058] Figure 5 Show Figure 4 A schematic diagram showing the separated state of the airfoil module and platform module of the stator blade assembly.

[0059] Figure 6 A perspective view showing the basic structure of the stator blade assembly according to a third embodiment of the present invention;

[0060] Figure 7 Show Figure 6 A schematic diagram showing the separated state of the airfoil module and platform module of the stator blade assembly.

[0061] Figure 8 A perspective view showing a first example of the mating surface of an airfoil module according to an embodiment of the present invention, wherein the microgroove structure is an array of multiple microgrooves;

[0062] Figure 9 A perspective view showing a second example of the mating surface of an airfoil module according to an embodiment of the present invention, wherein the microgroove structure is an annular groove;

[0063] Figure 10 A perspective view showing a third example of the mating surface of an airfoil module according to an embodiment of the present invention, wherein the microgroove structure is an arc-shaped groove;

[0064] Figure 11 A perspective view showing a fourth example of the mating surface of an airfoil module according to an embodiment of the present invention, wherein an annular groove is formed on the sidewall;

[0065] Figure 12 This is a partially enlarged cross-sectional schematic diagram showing the airfoil module and platform module of the stator blade assembly assembled together;

[0066] Figure 13 A schematic flowchart illustrating a method for manufacturing a turbine blade assembly according to an embodiment of the present invention is shown. Detailed Implementation

[0067] The invention will now be described in more detail with reference to the accompanying drawings. In the various drawings, the same elements are indicated by similar reference numerals. For clarity, the various parts in the drawings are not drawn to scale. Furthermore, some well-known parts may not be shown.

[0068] Many specific details of the invention, such as the structure, materials, dimensions, processing methods, and techniques of the devices, are described below to provide a clearer understanding of the invention. However, as those skilled in the art will understand, the invention may be implemented without following these specific details.

[0069] This invention can be presented in various forms, some of which will be described below.

[0070] Figure 1 A schematic cross-sectional view of the basic structure of the turbine engine involved in this embodiment is shown.

[0071] like Figure 1 As shown, the turbine engine as a whole includes, from front to back, the following components along the engine axis: an intake fan 110, a low-pressure compressor 120, a high-pressure compressor 130, a combustion chamber 140, a high-pressure turbine 150, a low-pressure turbine 160, and an exhaust cone 170.

[0072] At the front of the engine, the intake fan 110 is used to draw in outside air and perform primary pressurization on the airflow. The airflow entering the engine can be divided into two parts. One part enters the core engine flow path and flows sequentially through the low-pressure compressor 120, the high-pressure compressor 130, the combustion chamber 140, and the turbine section; the other part of the airflow flows along the outer bypass duct and is discharged from the rear of the engine to generate thrust.

[0073] Along the airflow direction, the low-pressure compressor 120 and the high-pressure compressor 130 constitute the engine's compression system. The low-pressure compressor 120 performs primary compression on the air entering the core engine, increasing its pressure and temperature; subsequently, the air enters the high-pressure compressor 130 for further compression, significantly increasing the air pressure. The compressed high-pressure air then enters the combustion chamber 140, where it mixes with fuel and burns, forming high-temperature, high-pressure combustion gas.

[0074] The high-temperature, high-pressure gas enters the turbine section axially and expands to perform work. First, the gas flows through the high-pressure turbine 150 and drives it to rotate. The high-pressure turbine 150 is driven by the high-pressure compressor 130 via a high-pressure rotor shaft, thus driving the high-pressure compressor 130. Subsequently, the gas continues to flow backward and drives the low-pressure turbine 160 to rotate. The low-pressure turbine 160 is driven by the low-pressure compressor 120 and the intake fan 110 via a low-pressure rotor shaft, thus driving the low-pressure compressor 120 and the intake fan 110. After performing work in the turbine section, the gas is finally discharged from the engine through the exhaust cone 170, thereby generating reaction thrust.

[0075] Structurally, the rotor blades and stator blades are mainly arranged in the internal flow channels of the low-pressure compressor 120, the high-pressure compressor 130, the high-pressure turbine 150, and the low-pressure turbine 160. For simplicity, Figure 1 Only a portion of the rotor blades 11 in the internal flow channel is shown. It can be understood that stator blades arranged in conjunction with the rotor blades 11 are also provided in the corresponding internal flow channel.

[0076] Specifically, in the compressor section, each stage typically includes a row of rotor blades arranged circumferentially and a row of stator blades located downstream of it. The rotor blades are mounted on the outer edge of the rotor disk and rotate at high speed with the rotor, doing work on the flowing air and increasing the total energy of the airflow; the stator blades are fixed to the inner wall of the casing, diffuse and guide the airflow, converting some kinetic energy into static pressure, and adjusting the airflow direction so that it enters the next stage rotor blades at a predetermined incident angle.

[0077] In the turbine section, each stage also includes a row of stator blades arranged circumferentially and a row of rotor blades located downstream of it. The stator blades, usually called nozzle guide vanes, are fixed to the turbine casing or bulkhead and are used to expand and accelerate the gas, converting the pressure energy of the gas into high-speed kinetic energy and guiding the airflow to the rotor blades. The rotor blades are mounted on the turbine disk and rotate under the action of high-speed airflow, converting the kinetic energy of the airflow into mechanical energy, which is then transmitted forward through the rotor shaft to drive the compressor and fan.

[0078] The aforementioned structure, in which rotor blades and stator blades are alternately arranged along the axial direction, constitutes the basic unit for energy conversion inside the engine, realizing the step-by-step conversion of combustion energy into mechanical energy.

[0079] Figure 2 A perspective view of a rotor blade assembly according to a first embodiment of the present invention is shown. The rotor blade assembly 10 includes an airfoil module 11 and a platform module 12. Figure 3 Show Figure 2 The diagram shows the separated state of the airfoil module 11 and platform module 12 of the rotor blade assembly.

[0080] The airfoil module 11 includes an airfoil body 111 extending along the spanwise direction of the turbine engine. The outer side of the airfoil body 111 has an external curved surface conforming to aerodynamic design, while the interior is provided with a complex network of cooling channels. At the bottom end of the airfoil module 11, i.e., the end near the platform module 12, a flange 112 extending laterally relative to the spanwise direction is formed. In embodiments, the airfoil module 11 is, for example, composed of a nickel-based superalloy (such as Inconel 718, René N5, or Mar-M247). The airfoil body 111 and the flange 112 of the airfoil module 11 are integrally formed metal structural parts using an additive manufacturing process. This additive manufacturing process includes at least one of selective laser melting, electron beam melting, directional energy deposition, or fused deposition modeling.

[0081] In this embodiment, selective laser melting is preferably employed. High-purity, uniformly sized pre-alloyed nickel-based superalloy powder is used as raw material. In a high-purity inert gas atmosphere, a high-energy-density laser beam scans and melts the powder layer by layer, precisely forming an integrated structure that includes complex internal cooling channels, thin-walled aerodynamic surfaces, and high-strength connecting flanges. After forming, the component undergoes post-processing including stress annealing, hot isostatic pressing to eliminate internal defects, and specific solution treatment and aging heat treatment for the superalloy to obtain a final component that meets the requirements for high-temperature strength, creep resistance, and fatigue resistance.

[0082] Platform module 12 is the basic structure that supports and carries airfoil module 11. A first mating surface is formed on the surface of flange 112 of airfoil module 11, and a second mating surface is formed on the surface of platform module 12; the two are fixed together by brazing. A tenon structure 121 is formed at the lower end of platform module 12 for insertion into turbine disk tenons to achieve circumferential positioning and torque transmission. Although not shown in the figures, the sides of platform module 12 are provided with guide holes and sealing grooves to collaboratively control the cooling airflow path and suppress high-temperature combustion gas leakage. Platform module 12 is, for example, composed of a nickel-based superalloy (such as Inconel 718, René N5, or Mar-M247) and manufactured using conventional manufacturing processes. These conventional manufacturing processes include at least one forming process selected from casting, forging, powder metallurgy, or precision casting, as well as machining processes.

[0083] In this embodiment, the platform module 12 blank is preferably manufactured using investment casting, and the material is a nickel-based high-temperature alloy that matches the airfoil module 11. After the cast blank undergoes hot isostatic pressing to eliminate casting defects such as internal shrinkage, it is precisely machined using precision machining processes such as CNC milling and electrical discharge machining to accurately machine the second mating surface for engaging with the flange 112, the precise profile of the tenon structure 121, and features such as guide holes and sealing grooves, ensuring that its dimensional tolerances and surface finish meet assembly and functional requirements.

[0084] During assembly, the first mating surface (flange 112 surface) of the airfoil module 11 and the second mating surface of the platform module 12 are precision ground and cleaned to ensure a smooth and clean mating surface. Then, a high-temperature nickel-based brazing filler metal is pre-placed or coated on one of the mating surfaces, and the two components are precisely aligned and fixed in a specialized fixture. The assembly is placed in a vacuum or protective atmosphere brazing furnace and heated according to a specific brazing temperature profile, causing the brazing filler metal to melt and fill the gaps between the mating surfaces under capillary action, achieving a strong metallurgical bond. After brazing, necessary stress-relief annealing and non-destructive testing (such as radiographic or ultrasonic testing) of the joint area can be performed to ensure the connection quality.

[0085] In the rotor blade assembly 10 of the first embodiment, the flange 112 in the airfoil module 11 is connected to the end of the airfoil body 111. The profile of the flange 112 in the airfoil module 11 is designed to match the profile shape of the platform module 12. The flange is used to expand the joint area between the airfoil module 11 and the platform module 12, thereby improving the joint strength and the uniformity of thermal stress distribution.

[0086] Figure 4 A perspective view showing the basic structure of a stator blade assembly according to a second embodiment of the present invention is shown. The stator blade assembly 20 includes an airfoil module 21, an outer platform module 22, and an inner platform module 23. Figure 5 Show Figure 4 The diagram shows the separated state of the airfoil module 21, outer platform module 22, and inner platform module 23 of the stator blade assembly.

[0087] The stator blade assembly 20 has an overall fan-shaped shape. Each stator blade assembly 20 includes two or more airfoil modules 21 arranged circumferentially. Each airfoil module 21 includes a radially extending airfoil body 211 and an outer platform module 22 and an inner platform module 23 respectively disposed on the radially outer and radially inner sides of the airfoil body 211, thus forming a single fan-shaped structure. Multiple stator blade assemblies 20 are spliced ​​circumferentially to form a stator ring structure. The outer platform modules 22 of the multiple stator blade assemblies 20 are mounted on the outer casing or guide housing, thereby positioning the stator blade assemblies between adjacent stages. The multiple inner platform modules 23, after circumferential splicing, form a continuous inner flow channel boundary structure.

[0088] In a turbine engine, the airflow generally flows from front to rear along the engine axis. High-temperature, high-pressure gas from upstream components first enters the annular airflow channel formed by the stator blade assembly 20 and flows between multiple airfoil modules 21. A convergent-throat-expanding flow channel is defined between two adjacent airfoil modules 21 to accelerate the airflow and change its direction, thereby providing the downstream rotor blades with airflow conditions that meet the design incident angle requirements.

[0089] In the stator blade assembly 20, the airfoil module 21 includes an airfoil body 211 extending along the spanwise direction of the turbine engine. The outer side of the airfoil body 211 forms an external curved surface conforming to aerodynamic design, and its interior also has a complex network of cooling channels. Compared with the airfoil module 11 of the rotor blade in the first embodiment, the airfoil module 21 of the stator blade has significant differences in aerodynamic design: its leading edge radius of curvature is generally smaller to optimize the guidance and segmentation of the incoming flow; its trailing edge is generally thinner to reduce wake loss; and its camber and angle of attack are designed primarily to guide the airflow to the optimal angle of attack of the downstream rotor blade. At both ends of the airfoil module 21, i.e., near the outer platform module 22 and the inner platform module 23, flanges 212 and 213 extending laterally relative to the spanwise direction are formed for connection with the corresponding platform modules.

[0090] In this embodiment, the airfoil module 21 is composed, for example, of a nickel-based superalloy (such as Inconel 718, René N5, or Mar-M247). The airfoil body 211 and flanges 212 and 213 of the airfoil module 21 are integrally formed metal structural parts using an additive manufacturing process. This additive manufacturing process includes at least one of selective laser melting, electron beam melting, directional energy deposition, or fused deposition modeling. In this embodiment, selective laser melting is also preferred. Using high-purity, uniformly distributed pre-alloyed nickel-based superalloy powder as raw material, the powder is melted layer by layer by a high-energy-density laser beam in a high-purity inert gas atmosphere to precisely form an integral structure containing complex internal irregular cooling channels, thin-walled aerodynamic surfaces, and high-strength connecting flanges at both ends. After forming, the component needs to undergo post-processing including stress annealing, hot isostatic pressing to eliminate internal defects, and specific solution and aging heat treatments for the superalloy to obtain a final component that meets the requirements for high-temperature strength, creep resistance, and fatigue resistance.

[0091] The outer platform module 22 and the inner platform module 23 form the basic structure supporting the airfoil module 21 and jointly define the radial boundary of the gas flow channel. A first outer mating surface is formed on the surface of the flange 212 of the airfoil module 21, and a second outer mating surface is formed on the surface of the outer platform module 22; the two are fixed together by brazing. A first inner mating surface is formed on the surface of the flange 213 of the airfoil module 21, and a second inner mating surface is formed on the surface of the inner platform module 23; the two are fixed together by brazing. The outer platform module 22 and the inner platform module 23 are, for example, made of nickel-based superalloys (such as Inconel 718, René N5, or Mar-M247) and manufactured using conventional manufacturing processes. These conventional manufacturing processes include at least one forming process selected from casting, forging, powder metallurgy, or precision casting, as well as machining processes.

[0092] In this embodiment, the blanks for the outer platform module 22 and the inner platform module 23 are preferably manufactured using investment casting, and the material is a nickel-based high-temperature alloy that matches the airfoil module 21. After the cast blanks are subjected to hot isostatic pressing to eliminate casting defects such as internal shrinkage, they are precisely machined using precision machining processes such as CNC milling and electrical discharge machining to precisely machine the second mating surface for engaging with the flanges 212 and 213 of the airfoil module 21, as well as features such as guide holes and sealing grooves, ensuring that their dimensional tolerances and surface finish meet the assembly and functional requirements.

[0093] During assembly, the first mating surfaces of the airfoil module 21 (the surfaces of flange 212 and flange 213) and the second mating surfaces of the outer platform module 22 and the inner platform module 23 are precision ground and cleaned to ensure that the mating surfaces are flat and clean. Then, a high-temperature nickel-based brazing filler metal is pre-placed or coated on one of the mating surfaces, and the three components are precisely aligned and fixed in a dedicated fixture. The assembly is placed in a vacuum or protective atmosphere brazing furnace and heated according to a specific brazing temperature profile, causing the brazing filler metal to melt and fill the gaps between the mating surfaces under capillary action, achieving a strong metallurgical bond. After brazing, necessary stress-relief annealing and non-destructive testing (such as radiographic or ultrasonic testing) of the joint area can be performed to ensure the connection quality.

[0094] In the rotor blade assembly 20 of the first embodiment, the flanges 212 and 213 in the airfoil module 21 are connected to the end of the airfoil body 211. The profile of the flange 212 in the airfoil module 21 is designed to match the profile shape of the outer platform module 22, and the profile of the flange 213 in the airfoil module 21 is designed to match the profile shape of the inner platform module 23. The flanges are used to increase the joint surface area between the airfoil module 21 and the outer platform module 22, as well as the joint surface area between the airfoil module 21 and the inner platform module 23, thereby improving the joint strength and the uniformity of thermal stress distribution.

[0095] Figure 6 A perspective view showing the basic structure of a stator blade assembly according to a third embodiment of the present invention is shown. The stator blade assembly 30 includes an airfoil module 31, an outer platform module 32, and an inner platform module 33. Figure 7 Show Figure 6 The diagram shows the separated state of the airfoil module 31, outer platform module 32, and inner platform module 33 of the stator blade assembly.

[0096] The stator blade assembly 30 has an overall fan-shaped shape. For simplicity and clarity, a simplified drawing of the stator blade assembly 30 is used as an example for illustration below. In the simplified drawing, the overall shape of the stator blade assembly 30 is shown as a linear segment, and a single airfoil module 31 is shown in a single stator blade assembly 30. It can be understood that in practical applications, the overall shape of the stator blade assembly 30 is still a fan-shaped segment that matches the shape of the casing, and the stator blade assembly 30 may include multiple airfoil modules 31 that share a common outer platform module 32 and an inner platform module 33.

[0097] In the stator blade assembly 30, the airfoil module 31 includes an airfoil body 311 extending along the spanwise direction of the turbine engine. The outer side of the airfoil body 311 forms an external curved surface conforming to aerodynamic design, and its interior also has a complex network of cooling channels. Compared with the airfoil module 11 of the rotor blade in the first embodiment, the airfoil module 31 of the stator blade has significant differences in aerodynamic design: its leading edge radius of curvature is generally smaller to optimize the guidance and segmentation of the incoming flow; its trailing edge is generally thinner to reduce wake loss; and its camber and angle of attack are designed primarily to guide the airflow to the optimal angle of attack of the downstream rotor blade. At both ends of the airfoil module 31, i.e., near the outer platform module 33 and the inner platform module 33, flanges 313 and 313 extending laterally relative to the spanwise direction are formed for connection with the corresponding platform modules.

[0098] The outer platform module 33 and the inner platform module 33 form the basic structure for supporting and carrying the airfoil module 31, and together define the radial boundary of the gas flow channel. A first outer mating surface is formed on the surface of the flange 312 of the airfoil module 31, and a second outer mating surface is formed in a groove on the surface of the outer platform module 32. The flange 312 of the airfoil module 31 is inserted into the groove on the surface of the outer platform module 32, and the two are fixed together by brazing. A first inner mating surface is formed on the surface of the flange 313 of the airfoil module 31, and a second inner mating surface is formed in a groove on the surface of the inner platform module 33. The flange 313 of the airfoil module 31 is inserted into the groove on the surface of the inner platform module 33, and the two are fixed together by brazing. The outer platform module 32 and the inner platform module 33 are, for example, made of nickel-based superalloys (such as Inconel 718, René N5, or Mar-M347) and manufactured using conventional manufacturing processes. These conventional manufacturing processes include at least one forming process selected from casting, forging, powder metallurgy, or precision casting, as well as machining processes.

[0099] In the rotor blade assembly 30 of the third embodiment, the flanges 312 and 313 in the airfoil module 31 are connected to the end of the airfoil body 311. The profile of the flange 312 in the airfoil module 31 is designed to conform to the end of the airfoil body 311, and the profile of the flange 313 in the airfoil module 31 is designed to conform to the end of the airfoil body 311. The flanges are used to expand the joint surface area between the airfoil module 31 and the outer platform module 32, as well as the joint surface area between the airfoil module 31 and the inner platform module 33, thereby improving the joint strength and the uniformity of thermal stress distribution.

[0100] Figure 8 A perspective view showing a first example of the mating surface of an airfoil module according to an embodiment of the present invention.

[0101] Similar in structure to the corresponding module of the stator blade assembly described in the above embodiments, the airfoil module 41 includes an airfoil body 411 and flanges 412 and 413 located at both ends, the contours of which are designed to match the contour shape of the corresponding platform module. Further, in Figure 8 The diagram shows a more detailed structure of the airfoil module.

[0102] In the airfoil module 41, the outer side of the airfoil body 411 forms an external curved surface conforming to aerodynamic design, and a complex cooling channel network is arranged inside. The cooling channel network is connected to the inlet 401 on the surface of the flange 412, allowing cooling gas to flow into the inlet 401 on the surface of the flange 412 through the corresponding inlet on the outer platform module, and thus into the interior of the cooling channel network. Furthermore, the cooling channel network is connected to multiple cooling holes 402 on the outer curved surface of the airfoil body 411, and after being discharged through the cooling holes, an attached cooling gas film is formed along the outer surface of the blade, thereby providing thermal protection for the airfoil body 411.

[0103] Microgroove structures 403 surrounding the cooling channel inlet 401 and outer blocking structures 404 surrounding the microgroove structures 403 are formed on the surfaces of flanges 412 and 413. The following detailed description uses flange 412 as an example only.

[0104] The microgroove structure 403 is a microgroove array comprising multiple microgrooves evenly distributed in an array on the surface. The depth of the microgrooves is, for example, from 10 micrometers to 200 micrometers. During brazing, the microgroove structure 403 serves as a storage and capillary channel for the brazing filler metal, guiding the molten filler metal to spread and fill more evenly along the joint interface. This helps to form a dense, continuous, and uniformly thick brazed joint, reducing defects such as incomplete penetration and porosity, and improving the quality and strength of the connection.

[0105] The outer blocking structure 404 is an annular groove surrounding the microgroove structure 403. During brazing, it restricts the radial overflow of molten filler metal at high temperatures into undesirable areas outside the joint surface (such as the flange sidewall or the non-jointed surface of the platform module) to avoid reducing the surface flatness of the stator blade assembly and compromising its aerodynamic characteristics.

[0106] Optionally, an inner blocking structure (not shown) surrounding the cooling channel inlet 401 is also formed on the surfaces of flanges 412 and 413. The inner blocking structure is located between the cooling channel inlet 401 and the microgroove structure 403, and is, for example, an annular groove surrounding the cooling channel inlet 401. During brazing, this restricts the radial overflow of molten filler metal at high temperatures into the cooling channel inlet and the interior of the cooling channel network at the joint surface, thereby ensuring unobstructed flow of the internal cooling network and reliable cooling performance while achieving a strong connection.

[0107] Optionally, the annular grooves of the outer and inner blocking structures can be partially filled with a flow-blocking material (such as a high-temperature resistant ceramic paste or metal oxide). This flow-blocking material remains solid or has a high viscosity at the brazing temperature, which can further enhance the blocking effect on the solder flow and provide more precise solder control.

[0108] In the aforementioned airfoil module's joint surface design, the brazing filler metal distribution is optimized through an array of microgroove structures 403, and the flow range of the brazing filler metal is precisely controlled through an outer blocking structure 404 and an optional inner blocking structure. This design effectively improves the uniformity and integrity of the brazed joint when brazing with the platform module, thereby enhancing the connection strength, sealing performance, and long-term reliability between the airfoil module 41 and the platform module. The flange 412 itself increases the joint area, which is beneficial for a more uniform distribution of interfacial thermal stress and mitigates the risk of thermal mismatch and interfacial delamination under high-temperature conditions.

[0109] Figure 9 A perspective view showing a second example of the mating surface of an airfoil module according to an embodiment of the present invention.

[0110] Similar in structure to the corresponding module of the stator blade assembly described in the above embodiments, the airfoil module 51 includes an airfoil body 511 and flanges 512 and 513 located at both ends, the profiles of which are designed to conform to the end shape of the airfoil body 511. Further, in Figure 9 The diagram shows a more detailed structure of the airfoil module.

[0111] In the airfoil module 51, the outer side of the airfoil body 511 forms an external curved surface conforming to aerodynamic design, and a complex cooling channel network is arranged inside. The cooling channel network is connected to the inlet 501 on the surface of the flange 512, allowing cooling gas to flow into the inlet 501 on the surface of the flange 512 through the corresponding inlet on the outer platform module, and thus into the interior of the cooling channel network. Furthermore, the cooling channel network is connected to multiple cooling holes 502 on the outer curved surface of the airfoil body 511, and after being discharged through the cooling holes, an attached cooling gas film is formed along the outer surface of the blade, thereby providing thermal protection for the airfoil body 511.

[0112] A microgroove structure 503 surrounding the cooling channel inlet 501 and an outer blocking structure 504 surrounding the microgroove structure 503 are formed on the surfaces of flanges 512 and 513. The following detailed description uses flange 512 as an example only.

[0113] The microgroove structure 503 includes multiple nested annular grooves, which are composed of multiple concentric and continuous annular grooves. The depth of these annular grooves is, for example, 10 micrometers to 200 micrometers. Figure 8 Unlike the array-type microgrooves shown, the annular groove structure 503 in this example provides a continuous, annular capillary flow path for the brazing filler metal. During brazing, the molten filler metal can preferentially spread uniformly and radially inward and outward along these annular channels, thereby more effectively promoting the uniform distribution of the filler metal across the entire joint surface. This helps to form a uniform, defect-free annular brazing seam, further improving the sealing uniformity and overall strength of the joint interface.

[0114] The outer blocking structure 504 is an annular groove surrounding the microgroove structure 503. During brazing, it restricts the radial overflow of molten filler metal at high temperatures into undesirable areas outside the joint surface (such as the flange sidewall or the non-jointed surface of the platform module) to avoid reducing the surface flatness of the stator blade assembly and compromising its aerodynamic characteristics.

[0115] Optionally, an inner blocking structure (not shown) surrounding the cooling channel inlet 501 is also formed on the surfaces of flanges 512 and 513. The inner blocking structure is located between the cooling channel inlet 501 and the microgroove structure 503, and is, for example, an annular groove surrounding the cooling channel inlet 501. During brazing, this restricts the radial overflow of molten solder into the cooling channel inlet and the interior of the cooling channel network at high temperatures, thereby ensuring unobstructed flow of the internal cooling network and reliable cooling performance while achieving a strong connection.

[0116] Optionally, the annular grooves of the outer and inner blocking structures can be partially filled with a flow-blocking material (such as a high-temperature resistant ceramic paste or metal oxide). This flow-blocking material remains solid or has a high viscosity at the brazing temperature, which can further enhance the blocking effect on the solder flow and provide more precise solder control.

[0117] Optionally, the sidewalls 505 of flanges 512 and 513 also serve as additional mating surfaces. When the flanges 512 and 513 of the airfoil module 51 are inserted into the surface grooves of the corresponding platform module, a circumferential mating gap is formed between them and the sidewalls of the surface grooves of the corresponding platform module. During brazing, brazing filler metal is accommodated in the circumferential mating gap, further enhancing the circumferential connection strength and sealing between the flanges of the airfoil module 51 and the grooves of the corresponding platform module.

[0118] In the aforementioned airfoil module's joint surface design, a microgroove structure 503 composed of nested annular grooves optimizes the storage and capillary flow path of the brazing filler metal, resulting in a more uniform and controllable distribution. The flow range of the brazing filler metal is precisely controlled through an outer blocking structure 504 and an optional inner blocking structure. This design effectively improves the uniformity and integrity of the brazed joint when brazing with the platform module, thereby enhancing the connection strength, sealing performance, and long-term reliability between the airfoil module 51 and the platform module. Flanges 512 and 513 themselves increase the joint area, facilitating a more uniform distribution of interfacial thermal stress and mitigating the risks of thermal mismatch and interfacial delamination under high-temperature conditions.

[0119] Figure 10 A perspective view showing a third example of the mating surface of an airfoil module according to an embodiment of the present invention.

[0120] Similar in structure to the corresponding module of the stator blade assembly described in the above embodiments, the airfoil module 61 includes an airfoil body 611 and flanges 612 and 613 located at both ends, the profiles of which are designed to conform to the end shape of the airfoil body 611. Further, in Figure 10 The diagram shows a more detailed structure of the airfoil module.

[0121] In the airfoil module 61, the outer side of the airfoil body 611 forms an external curved surface conforming to aerodynamic design, and a complex cooling channel network is arranged inside. The cooling channel network is connected to the inlet 601 on the surface of the flange 612, allowing cooling gas to flow into the inlet 601 on the surface of the flange 612 through the corresponding inlet on the outer platform module, and thus into the interior of the cooling channel network. Furthermore, the cooling channel network is connected to multiple cooling holes 602 on the outer curved surface of the airfoil body 611, and after being discharged through the cooling holes, an attached cooling gas film is formed along the outer surface of the blade, thereby providing thermal protection for the airfoil body 611.

[0122] A microgroove structure 603 surrounding the cooling channel inlet 601 and an outer blocking structure 604 surrounding the microgroove structure 603 are formed on the surfaces of flanges 612 and 613. The following detailed description uses flange 612 as an example only.

[0123] The airfoil module 61 of the third example is substantially the same in structure as the airfoil module 51 of the second example. Optionally, an inner blocking structure (not shown) surrounding the cooling channel inlet 601 is also formed on the surfaces of flanges 612 and 613. Optionally, the sidewalls 605 of flanges 612 and 613 also serve as additional mating surfaces. The difference in the airfoil module 61 of the third example lies in the different microgroove structure design of the mating surfaces.

[0124] In the airfoil module 61, the microgroove structure 603 includes multiple nested arc-shaped grooves, that is, it is composed of multiple discontinuous, arc-shaped nested microgrooves. These arc-shaped microgrooves are distributed circumferentially at intervals and form multiple nested but discontinuous annular patterns around the inlet 601. The depth of each arc-shaped microgroove is, for example, 10 micrometers to 200 micrometers.

[0125] and Figure 9 Compared to the microgroove structure shown, the microgroove structure 603 employs nested arc-shaped grooves, providing discrete and controlled solder storage points during brazing. This allows for more flexible adaptation to minor fluctuations in local gaps or surface unevenness at the joint surface. Because the solder is released from multiple discrete point sources, rather than relying on a continuous annular capillary channel, the requirement for uniformity in the assembly gap is relatively reduced. This helps to form a uniform and continuous brazed joint within a wider process window, reducing the risk of solder loss due to excessively large local gaps or insufficient solder filling due to excessively small local gaps.

[0126] Figure 11 A perspective view showing a fourth example of the mating surface of an airfoil module according to an embodiment of the present invention.

[0127] Similar in structure to the corresponding module of the stator blade assembly described in the above embodiments, the airfoil module 71 includes an airfoil body 711 and flanges 712 and 713 located at both ends, the profiles of which are designed to conform to the end shape of the airfoil body 711. Further, in Figure 11 The diagram shows a more detailed structure of the airfoil module.

[0128] In the airfoil module 71, the outer side of the airfoil body 711 forms an external curved surface conforming to aerodynamic design, and a complex cooling channel network is arranged inside. The cooling channel network is connected to the inlet 701 on the surface of the flange 712, allowing cooling gas to flow into the inlet 701 on the surface of the flange 712 through the corresponding inlet on the outer platform module, and thus into the interior of the cooling channel network. Furthermore, the cooling channel network is connected to multiple cooling holes 702 on the outer curved surface of the airfoil body 711, and after being discharged through the cooling holes, an attached cooling gas film is formed along the outer surface of the blade, thereby providing thermal protection for the airfoil body 711.

[0129] A microgroove structure 703 surrounding the cooling channel inlet 701 and an outer blocking structure 704 surrounding the microgroove structure 703 are formed on the surfaces of flanges 712 and 713. The following description uses flange 712 as an example.

[0130] The airfoil module 71 of the fourth example is substantially the same in structure as the airfoil module 61 of the third example. Optionally, an inner blocking structure (not shown) surrounding the cooling channel inlet 701 is also formed on the surfaces of the flanges 712 and 713. The difference of the airfoil module 71 of the fourth example is that additional microgrooving structures 706 are provided on the sidewalls 705 of the flanges 712 and 713. The microgrooving structure 706 is, for example, an annular groove formed on the sidewall.

[0131] and Figure 10 Compared to the segmented arc-shaped microgroove structure shown, microgroove structure 706 is an additional microgroove structure set on the flange sidewall. This design extends solder control from the main mating surface to the three-dimensional mating interface. By providing dedicated solder storage and flow channels for the lateral mating surfaces, it ensures a dense and continuous metallurgical bond between the flange and the platform module in the lateral direction. This not only significantly enhances the mechanical strength and overall sealing performance of the connection structure in the three-dimensional direction, making it particularly suitable for resisting complex loads generated by centrifugal force, aerodynamic force, and thermal stress, but also optimizes the filling reliability of the three-dimensional gap in the flange embedded assembly, reducing the dependence on solder flow on a single main mating surface, thereby greatly improving the long-term reliability and process robustness of the brazed joint under harsh service conditions.

[0132] Figure 12 This diagram shows a partially enlarged cross-sectional view of the stator blade assembly, with the airfoil module and platform module assembled together. The stator blade assembly 30 includes an airfoil module 31, an outer platform module 32, and an inner platform module 33. Figure 12 Only a portion of the structure of the airfoil module 31 and the outer platform module 32 is shown, and a cooling hole 302 communicating with the cooling channel inlet 301 is also shown.

[0133] Cooling channel inlet 301 is formed on the bottom surface of flange 312 and communicates with the cooling channel network inside airfoil module 31. Cooling hole 302 is opened on the outer curved surface of airfoil body 311 and is connected to the internal cooling channel network. In the assembled state, cooling gas (indicated by arrow) from the upstream air source of outer platform module 32 enters the internal cooling network through inlet 301, flows through the complex channel to convect and cool the metal substrate, and finally exits from cooling hole 302, forming a protective gas film on the outer surface of the blade.

[0134] As shown in the figure, the flange 312 of the airfoil module 31 and the recessed surface (i.e., the second mating surface) of the outer platform module 32 are metallurgically bonded through brazing. A microgroove structure 303 is provided on the surface of the flange 312 (the first mating surface), and this microgroove structure 303 can be as described above. Figure 8 to Figure 11 The array, annular groove, arc-shaped groove, or combination thereof are shown. During brazing, the pre-placed filler metal 321 melts at high temperature. The microgroove structure 303 forms spatial concave-convex geometric features in the interface region, which can serve as a channel for receiving and guiding the filler metal, guiding the filler metal to spread and fill more evenly along the joint interface, forming a continuous metallurgical bonding layer in multiple directions. This helps to form a dense, continuous, and uniformly thick brazed seam, reducing the generation of defects such as incomplete penetration and porosity, and further ensuring the connection quality.

[0135] Preferably, the microgroove structure 303 is positioned away from the peak region of the main load-bearing stress of the airfoil module 31 and / or the abrupt change region of the thermal gradient. In the figure, the stress distribution curve shows that the peak region of the main load-bearing stress of the airfoil module 31 is located at the blade root. Therefore, the microgroove structure 303 is located away from the high-stress concentration region at the blade root. Furthermore, the microgroove structure 303 is also located away from the high-temperature and high-temperature gradient regions of the aerodynamic outer surface, ensuring that the brazed interface does not overlap with the peak of the main load-bearing stress (and / or the abrupt change region of the thermal gradient), thereby helping to improve the stability and durability of the interface under high-temperature thermo-mechanical cyclic loading.

[0136] Figure 13 A schematic flowchart illustrating a method for manufacturing a turbine blade assembly according to an embodiment of the present invention is shown. The manufacturing method includes steps S01 to S03.

[0137] In step S01, at least one airfoil module is formed using an additive manufacturing process.

[0138] Specifically, the airfoil module includes: an airfoil body extending along the spanwise direction of the turbine engine, including an outer curved surface formed on its outer side and a cooling channel network disposed therein; and a flange located at the end of the airfoil body, extending laterally relative to the spanwise direction, thereby forming a first mating surface on the surface of the flange.

[0139] The airfoil body and the flange are integrally formed using an additive manufacturing process. The additive manufacturing process includes at least one of selective laser melting, electron beam melting, directional energy deposition, or fused deposition modeling. Preferably, selective laser melting is used, with nickel-based superalloy powder as raw material, and the mixture is melted layer by layer under inert gas protection to obtain an integrated structure containing complex internal cooling channels, external aerodynamic surfaces, and a transverse flange. After forming, the at least one airfoil module undergoes post-processing, including stress annealing, hot isostatic pressing to eliminate internal defects, and solution and aging heat treatment tailored to the material properties, to obtain a final component that meets high-temperature performance requirements.

[0140] Optionally, a first microgroove structure for receiving solder is formed on the surface of the flange. This first microgroove structure may be an array of nested annular grooves or nested arcuate grooves, with a depth of, for example, 10 micrometers to 200 micrometers. Further optionally, an outer blocking structure (e.g., annular groove) is formed around the first microgroove structure on the surface of the flange, and / or an inner blocking structure (e.g., annular groove) is formed around the inlets and / or outlets of the cooling channel network formed on the flange surface. The annular grooves of the blocking structure may be partially filled with a flow-blocking material to enhance the flow-blocking effect. Additionally, when the flange is designed as a recessed surface embedded in a platform module, a second microgroove structure for receiving solder may be formed on the sidewall surface of the flange.

[0141] In step S02, the platform module is formed using a conventional manufacturing process.

[0142] Specifically, the platform module is provided with a second mating surface that matches the first mating surface of the at least one airfoil module. The second mating surface can be a flat surface or a recessed surface of the platform module for planar mating or embedded mating with the flange of the airfoil module.

[0143] The platform module is made of, for example, a nickel-based superalloy and manufactured using conventional manufacturing processes. These conventional processes include at least one forming process selected from casting, forging, powder metallurgy, or precision casting, followed by subsequent machining. Preferably, a blank is manufactured using investment casting, and after hot isostatic pressing, the second mating surface, necessary mechanical positioning structures (such as dovetail joints, tongue joints, groove joints, keyways, etc.), guide holes, and sealing grooves are precisely formed using CNC machining. For the stator blade assembly, the platform module can be an outer platform module and / or an inner platform module, and can be designed as a fan-shaped segment to connect with multiple airfoil modules.

[0144] In step S03, a brazing process is used to join the at least one airfoil module to the platform module.

[0145] First, the first mating surface of the airfoil module and the second mating surface of the platform module are precisely cleaned and prepared.

[0146] Subsequently, brazing filler metal is pre-applied or coated on one or both mating surfaces. The brazing filler metal is preferably a nickel-based low-melting-point filler metal containing nickel, chromium, boron, and silicon. The airfoil module and platform module are precisely aligned and fixed in a dedicated fixture, ensuring that the assembly gap between the flange and the second mating surface meets brazing requirements. The assembly is placed in a vacuum or protective atmosphere furnace and heated according to a set brazing temperature profile. During heating, the brazing filler metal melts and, under capillary action, is guided by the first microgroove structure (and optional first microgroove structure) on the flange surface, uniformly distributing and filling the mating interface between the airfoil module and the platform module. Simultaneously, the outer blocking structure restricts the radial overflow of the brazing filler metal to the periphery of the mating surface during high-temperature melting, and the inner blocking structure prevents the brazing filler metal from covering the inlet and / or outlet of the cooling channel. After solidification, the brazing filler metal forms a dense, continuous metallurgical bond, firmly connecting the two components.

[0147] After brazing, stress-relief annealing can be performed, and non-destructive testing can be conducted on the joint area to ensure connection quality.

[0148] In the above manufacturing method, the flange of the at least one airfoil module expands the area of ​​the first mating surface, and after being joined with the platform module, it significantly improves the joint strength between the modules and the uniformity of thermal stress distribution.

[0149] In this document, relational terms such as "first" and "second" are used merely to distinguish one entity or operation from another, without necessarily requiring or implying any such actual relationship or order between these entities or operations. Furthermore, the terms "comprising," "including," or any other variations thereof are intended to cover non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements includes not only those elements but also other elements not expressly listed, or elements inherent to such a process, method, article, or apparatus. Without further limitation, an element defined by the phrase "comprising one..." does not exclude the presence of other identical elements in the process, method, article, or apparatus that includes said element.

[0150] As described above, these embodiments of the present invention do not exhaustively cover all details, nor do they limit the invention to the specific embodiments described. Clearly, many modifications and variations can be made based on the above description. This specification selects and specifically describes these embodiments to better explain the principles and practical applications of the invention, thereby enabling those skilled in the art to effectively utilize the invention and its modifications. The invention is limited only by the claims and their full scope and equivalents.

Claims

1. An airfoil module for a turbine engine, comprising: The airfoil body extends along the spanwise direction of the turbine engine and includes an outer curved surface formed on its outer side and a network of cooling channels disposed inside it. as well as A flange, located at the end of the airfoil body, extends laterally relative to the spanwise direction, thereby forming a first mating surface on the surface of the flange.

2. The airfoil module according to claim 1, wherein, The outer curved surface of the airfoil body has multiple cooling holes that communicate with the cooling channel network.

3. The airfoil module according to claim 1, wherein, The surface of the flange is formed with a first microgroove structure for receiving brazing filler metal.

4. The airfoil module according to claim 3, wherein, The depth of the first microgroove structure is 10 micrometers to 200 micrometers.

5. The airfoil module according to claim 3, wherein, The position of the first microgroove structure is offset from the steep transition zone of the main load-bearing stress peak region and / or thermal gradient region of the airfoil module.

6. The airfoil module according to claim 5, wherein, The first microgroove structure is located away from the blade root of the airfoil module.

7. The airfoil module according to claim 3, wherein, The first microgroove structure includes multiple microgrooves arranged in an array.

8. The airfoil module according to claim 3, wherein, The first microgroove structure includes multiple nested annular grooves.

9. The airfoil module according to claim 3, wherein, The first microgroove structure includes multiple nested arc-shaped grooves.

10. The airfoil module according to claim 3, wherein, An outer and / or inner blocking structure adjacent to the first microgroove structure is formed on the surface of the flange to limit radial overflow of the solder during the high-temperature melting process.

11. The airfoil module according to claim 10, wherein, The outer blocking structure and the inner blocking structure both include an annular groove.

12. The airfoil module according to claim 11, wherein, The outer blocking structure and / or the inner blocking structure also include flow-blocking material that partially fills the annular groove.

13. The airfoil module according to claim 1, wherein, The airfoil body and the flange are integrally formed using additive manufacturing process.

14. The airfoil module according to claim 1, wherein, The flange has a contour shape that is conformal or non-conformal to the end of the airfoil body.

15. The airfoil module according to claim 14, wherein, The flange sidewall surface is formed with a second microgroove structure for accommodating brazing filler metal.

16. A turbine blade assembly, comprising: At least one airfoil module according to any one of claims 1 to 15; as well as A platform module, wherein the platform module is provided with a second mating surface that matches the first mating surface of the at least one airfoil module.

17. The turbine blade assembly of claim 16, wherein, The second mating surface of the platform module is a flat surface or a recessed surface of the platform module, and the flange of the airfoil module is embedded in the recessed surface of the platform module.

18. The turbine blade assembly of claim 16, wherein, The turbine blade assembly is a stator blade assembly, and the airfoil module of the stator blade assembly includes at least one airfoil body, a first flange located at the top of the at least one airfoil body, and a second flange located at the bottom of the at least one airfoil body.

19. The turbine blade assembly of claim 18, wherein, The stator blade assembly is a sector segment, and the at least one airfoil module includes multiple airfoil modules that share the platform module, and the multiple airfoil modules are distributed circumferentially along the sector segment.

20. A method for manufacturing a turbine blade assembly for a turbine engine, comprising: At least one airfoil module is formed using additive manufacturing processes; Platform modules are formed using traditional manufacturing processes; as well as The at least one airfoil module is joined to the platform module using a brazing process. The at least one airfoil module includes: an airfoil body extending along the spanwise direction of the turbine engine, including an outer curved surface formed on its outer side and a cooling channel network disposed therein; and a flange located at the end of the airfoil body, extending laterally relative to the spanwise direction, thereby forming a first mating surface on the surface of the flange. The platform module is provided with a second mating surface that matches the first mating surface of the at least one airfoil module. The flange of the at least one airfoil module increases the area of ​​the first mating surface to improve the joint strength and the uniformity of thermal stress distribution.

21. The manufacturing method according to claim 20, wherein, The surface of the flange is formed with a first microgroove structure for receiving brazing filler metal; and / or The sidewall surface of the flange is formed with a second microgroove structure for receiving the brazing filler metal. During the brazing process, the first microgroove structure and the second microgroove structure guide the brazing filler metal to be evenly distributed along the interface between the at least one airfoil module and the platform module.

22. The manufacturing method according to claim 21, wherein, The additive manufacturing process includes at least one of selective laser melting, electron beam melting, directional energy deposition, or fused deposition modeling.

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