Turbine multi-guide vane throat area control method, turbine and aero-engine

Abstract

This disclosure relates to a method for controlling the throat area of ​​a multi-stage guide vane of a turbine, a turbine, and an aero-engine. The control method includes: based on the design shape of a blade that meets aerodynamic performance requirements, offsetting the shape inward by a uniform first coating thickness to form a first theoretical blade model; after processing a first blade substrate according to the first theoretical blade model, connecting at least two first blade substrates into one unit, with a shielding area between adjacent first blade substrates; spraying a coating layer on at least two first blade substrates according to the first coating thickness, and measuring the actual thickness distribution of the coating; based on the design shape of a blade that meets aerodynamic performance requirements, offsetting the shape inward by the actual thickness distribution to form a second theoretical blade model; after processing a second blade substrate according to the second theoretical blade model, connecting at least two second blade substrates into one unit; and spraying a coating layer on at least two second blade substrates according to the first coating thickness to obtain a multi-stage blade product.

Description

Technical Field

[0001] This disclosure relates to the field of aero-engine technology, and in particular to a method for controlling the throat area of ​​a multi-stage guide vane of a turbine, a turbine, and an aero-engine. Background Technology

[0002] High-pressure turbine flow capacity and aerodynamic efficiency are key aerodynamic performance parameters for aero-engines. Designers continuously optimize blade and flow channel shapes to achieve higher aerodynamic efficiency, and control key dimensions during manufacturing to ensure blade profile and throat area, thus achieving design goals for aerodynamic efficiency and turbine flow capacity. Therefore, the physical shapes of high-pressure turbine blades and flow channels should be as close as possible to the design objectives; deviations in shape often lead to a decrease in turbine aerodynamic performance.

[0003] The first-stage guide vane of a high-pressure turbine is crucial to the turbine's aerodynamic efficiency, especially its flow capacity. Located at the combustion chamber outlet, this guide vane is subjected to the erosion of high-temperature combustion gases over a long period. To ensure the reliability and service life of the guide vane, an advanced and complex cooling design is required, and a heat-insulating coating is sprayed onto the surfaces of the guide vanes and their rims to reduce the metal wall temperature to an acceptable range.

[0004] The commonly used blade design method in current engineering is as follows: First, obtain the blade profile that meets the turbine's flow capacity and aerodynamic efficiency. Based on this profile, a certain design coating thickness is added to create a blade structure model, resulting in the profile of the casting. Therefore, once the casting is completed, only the designed coating thickness needs to be applied to obtain the target profile. Furthermore, the circumferential profiles of all blades in the entire ring guide assembly are identical, and the coating thickness on the blade back and blade base sides is the same.

[0005] The first-stage guide vanes of a high-pressure turbine form several sets of multi-section blades along the entire circumferential direction. This means that at least two adjacent guide vanes are first welded together and then coated to form a multi-section blade. Due to the narrow throat between adjacent blades, there are obstructed areas within the profile of the double or multi-section guide vane channel, resulting in uncoated or thinly coated areas. This adversely affects turbine performance, causing engine components to operate outside their design parameters, leading to reduced engine performance or service life. Summary of the Invention

[0006] This disclosure provides a method for controlling the throat area of ​​multi-stage guide vanes in a turbine, a turbine, and an aero-engine, which can ensure that the turbine has superior working performance.

[0007] The first aspect of this disclosure provides a method for controlling the throat area of ​​a multi-stage guide vane in a turbine. The turbine includes at least one set of blades arranged axially, each set of blades including multiple blades arranged circumferentially, and at least two adjacent blades circumferentially connected together before coating to form a multi-stage vane. The control method includes:

[0008] Based on the design blade shape that meets aerodynamic performance requirements, the shape is offset inward according to the uniform first coating thickness to form the first theoretical blade model.

[0009] After the first blade substrate is formed according to the first theoretical blade model, at least two first blade substrates are connected into one piece, and there is a shielding area between two adjacent first blade substrates, and the throat of the blade is located in the shielding area.

[0010] The coating is sprayed onto at least two first blade substrates according to the first coating thickness, and the actual thickness distribution of the coating is measured;

[0011] Using the actual thickness distribution of the coating as the input for bias, based on the design blade shape that meets aerodynamic performance requirements, the shape is biased inward according to the actual thickness distribution to form a second theoretical blade model.

[0012] After the second blade substrate is formed according to the second theoretical blade model, at least two second blade substrates are connected into one piece;

[0013] A coating is sprayed onto at least two second blade substrates according to the first coating thickness to obtain a multi-blade product.

[0014] In some embodiments, the step of designing the shape of the blades to meet aerodynamic performance requirements includes:

[0015] Select at least three radial height positions of the flow channel, covering at least the inner diameter, outer diameter, and middle diameter of the flow channel, and design two-dimensional airfoil profiles that meet aerodynamic performance, including the back of the blade and the blade base. The minimum distance between the back of the blade and the blade base of two adjacent two-dimensional airfoil profiles is the throat.

[0016] By adopting the radial stacking law, two-dimensional blade profiles with different radial heights of the flow channel are radially stacked to form a three-dimensional blade shape that meets aerodynamic performance requirements.

[0017] In some embodiments, the step of measuring the actual thickness distribution of the coating includes:

[0018] Take at least three cross sections along the blade height direction, covering the entire blade height;

[0019] Two guide blades are cut along each cross section to obtain a two-dimensional blade profile. The coating thickness distribution on the surface of the two-dimensional blade profile at each cross section is measured to obtain the actual coating spraying thickness distribution.

[0020] In some embodiments, in the actual thickness distribution of the coating, there is a second coating thickness between two adjacent first blade substrates in the shielded area and a first coating thickness in the unshielded area, wherein the second coating thickness is less than the first coating thickness.

[0021] In some embodiments, a two-dimensional leaf profile is formed in a cross section of the second blade substrate perpendicular to the blade height direction. The two-dimensional leaf profile includes a leaf back and a leaf base. The position on the leaf back corresponding to the throat is a second point. A first point and a third point are selected on both sides of the second point along the leaf back. The first point and the third point are located at the edges of the shading area. The coating thickness at the second point is zero, and the actual coating thickness gradually increases from the second point to the first point and from the second point to the third point to the first coating thickness.

[0022] In some embodiments, a two-dimensional leaf profile is formed in a cross section of the second blade substrate perpendicular to the blade height direction. The two-dimensional leaf profile includes a leaf back and a leaf base. The two ends of the shading area of ​​the leaf base are the fourth point and the fifth point, respectively. The fourth point is located at the tail edge. The position on the leaf base corresponding to the throat is located between the fourth point and the fifth point. The coating thickness at the fourth point is zero. The actual coating thickness gradually increases from the fourth point to the fifth point to the first coating thickness.

[0023] A second aspect of this disclosure provides a turbine, including at least one set of guide vanes arranged along an axial direction. Each set of guide vanes includes a plurality of vanes arranged circumferentially, and at least two adjacent vanes along the circumferential direction are connected to form a multi-stage vane. The multi-stage vane is formed according to the turbine multi-stage guide vane throat area control method of the above embodiments.

[0024] In some embodiments, the blades have rim plates, the rim plates of two adjacent blades in a multi-blade configuration are connected to each other and there is no gap at the connection interface, and the rim plates of two adjacent blades in a multi-blade configuration are connected to each other and there is a gap at the connection interface.

[0025] In some embodiments, the blade group is a guide blade group.

[0026] A third aspect of this disclosure provides an aircraft engine, including the turbine of the above embodiments.

[0027] The method for controlling the throat area of ​​multi-stage guide vanes in this disclosure measures the actual coating thickness distribution of each blade in the multi-stage guide vane as input for coating bias in blade structure modeling. This ensures that the actual shape of each blade after the second blade substrate is coated is the same as the aerodynamic design target shape. It prevents the absence of coating or thinning of coating in areas where the multi-stage blades need to be pre-connected before coating, making the actual shape of the blade more consistent with the actual design target. When the obstructed area is located at the throat, it can prevent the throat area from being too low, making the throat area more consistent with the design target, thereby ensuring the turbine's flow capacity and aerodynamic performance, and improving the performance and lifespan of the aero-engine. Attached Figure Description

[0028] To more clearly illustrate the technical solutions in the embodiments of this disclosure or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are only some embodiments of this disclosure. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0029] Figure 1 This is a schematic diagram showing the throat area design between adjacent first and second blades in the multi-blade configuration of this disclosure.

[0030] Figure 2 This is a schematic diagram of forming a three-dimensional airfoil by stacking two-dimensional airfoils.

[0031] Figure 3 This is the control line for the distribution of coating thickness on the blade back.

[0032] Figure 4 This is the control line for the distribution of coating thickness on the leaf plate.

[0033] Figure 5 This is the thickness distribution curve of the coating on the back of the blade.

[0034] Figure 6 This is the thickness distribution curve of the leaf-pot coating.

[0035] Figure 7 This is a schematic diagram of the installation of multi-blade units. Detailed Implementation

[0036] The technical solutions of the embodiments of this disclosure will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of this disclosure, and not all embodiments. The following description of at least one exemplary embodiment is merely illustrative and is in no way intended to limit this disclosure or its application or use. All other embodiments obtained by those skilled in the art based on the embodiments of this disclosure without creative effort are within the scope of protection of this disclosure.

[0037] Techniques, methods, and equipment known to those skilled in the art may not be discussed in detail, but where appropriate, they should be considered part of the specification.

[0038] In the description of this disclosure, it should be understood that the terms "center," "lateral," "longitudinal," "front," "rear," "left," "right," "upper," "lower," "vertical," "horizontal," "top," "bottom," "inner," and "outer," etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings, and are only for the convenience of describing this disclosure and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation, and therefore should not be construed as a limitation on the scope of protection of this disclosure.

[0039] In the description of this disclosure, it should be understood that the use of terms such as "first" and "second" to define components is merely for the purpose of distinguishing the corresponding components. Unless otherwise stated, the above terms have no special meaning and therefore should not be construed as limiting the scope of protection of this disclosure.

[0040] The inventors noted that having uncoated or thinly coated areas in multi-stage turbine blades can adversely affect turbine performance. The main reasons are as follows: First, uneven coating thickness on the guide vane surface makes the actual blade profile uneven and deviates from the design, especially for the profile near the throat where airflow velocity is high, which leads to increased aerodynamic losses and reduced efficiency. Second, uneven throat area in the circumferential blade channels of the entire ring guide vane increases aerodynamic mixing losses, and the throat area of ​​the entire ring guide vane deviates from the design. These factors can cause engine components to operate in an undesigned state, resulting in reduced engine performance or service life.

[0041] Therefore, this disclosure provides a method for controlling the throat area of ​​multi-stage turbine guide vanes, hereinafter referred to as the "throat area control method". For example... Figure 7As shown, the turbine includes at least one set of guide vanes arranged axially. Each set of guide vanes includes multiple blades arranged circumferentially, with an even number of blades. At least two adjacent blades circumferentially are joined together before painting to form a multi-blade assembly. For example, the connection can be achieved by welding, or more specifically, by precision casting. For example, each multi-blade assembly includes a first blade 1 and a second blade 2. The first blade 1 and the second blade 2 have the same profile and are painted after welding, with the first blade and the second blade 2 having the same coating thickness distribution. For example, the turbine is a high-pressure turbine.

[0042] Based on this type of guide vane assembly, in some embodiments, the throat area control method includes:

[0043] Step 110: Based on the design blade shape that meets aerodynamic performance requirements, offset the shape inward according to the uniform first coating thickness to form the first theoretical blade model.

[0044] Step 120: After forming the first blade substrate according to the first theoretical blade model, connect at least two first blade substrates into one piece. There is a shielding area between two adjacent first blade substrates, and the throat of the blade is located in the shielding area.

[0045] Step 130: Spray a coating onto at least two first blade substrates according to the first coating thickness, and measure the actual thickness distribution of the coating;

[0046] Step 140: Using the actual thickness distribution of the coating as the input for the offset, based on the design blade shape that meets aerodynamic performance requirements, the shape is offset inward according to the actual thickness distribution to form the second theoretical blade model.

[0047] Step 150: After forming the second blade substrate according to the second theoretical blade model, connect at least two second blade substrates into one piece;

[0048] Step 160: Apply a coating to at least two second blade substrates according to the first coating thickness to obtain a multi-blade product.

[0049] Steps 110-160 are executed sequentially.

[0050] In step 110, the aerodynamically designed blade refers to one that meets the turbine flow capacity and aerodynamic efficiency requirements. Based on the shape of this designed blade, a first theoretical blade model is formed by offsetting the first coating thickness T1 (e.g., 0.3 mm) inward along the entire circumference of the blade. In step 120, after forming the first blade substrate, for example by casting, at least two first blade substrates are welded together. For example, the first blade substrates of first blade 1 and second blade 2 are welded together, and the region of first blade 1 near the trailing edge and the region of second blade 2 near the leading edge have a shielding region. The minimum distance between the first blade 1 and the second blade 2 in the shielding region is the throat 7 of the blade.

[0051] In step 130, with a first coating thickness T1 as the target, a coating is sprayed onto at least two first blade substrates. On the surface of the unmasked area, it is easy to ensure the first coating thickness T1; however, in the masked area, the coating thickness is difficult to reach the first coating thickness T1. Therefore, the purpose of step 130 is to obtain the actual thickness distribution on the blade surface. In step 140, based on the design blade shape that meets aerodynamic performance requirements, a second round of blade design is performed. The shape is offset inward according to the actual thickness distribution with unequal thicknesses, forming a second theoretical blade model.

[0052] In step 150, the blade undergoes a second round of processing to form the second blade substrate. At least two second blade substrates are then welded together. In step 160, a first coating thickness T1 is applied. Although the coating thickness is thinner in the shaded area, the surface of the second blade substrate in the shaded area compensates for this reduction. Mass production is achieved under stable coating application conditions, thus ensuring the blade's shape and throat area remain consistent, and the throat area is the same at all circumferential positions of the entire guide ring.

[0053] This embodiment measures the actual coating thickness distribution of each blade in the multi-stage guide vane as input for the coating bias in blade structure modeling. This ensures that the actual shape of each blade after the second blade substrate is coated is the same as the aerodynamic design target shape. It prevents the absence of coating or thinning of coating in areas where the multi-stage blades need to be pre-connected before coating. This makes the actual shape of the blade more consistent with the actual design target. When the obstructed area is located at the throat, it can prevent the throat area from being too low, making the throat area more consistent with the design target. This ensures the turbine's flow capacity and aerodynamic performance, and improves the performance and lifespan of the aero-engine.

[0054] In some embodiments, such as Figure 2 As shown, in step 110, the step of designing the shape of the blade to meet aerodynamic performance includes:

[0055] Select at least three radial height positions of the flow channel, covering at least the inner diameter A, outer diameter C and middle diameter B of the flow channel, and design two-dimensional airfoil profiles that meet aerodynamic performance, including the back of the blade 3 and the blade face 4. The minimum distance between the back of the blade 3 and the blade face 4 of two adjacent two-dimensional airfoil profiles is the throat 7.

[0056] By adopting the radial stacking law, two-dimensional blade profiles with different radial heights of the flow channel are radially stacked to form a three-dimensional blade shape that meets aerodynamic performance requirements.

[0057] This embodiment forms a three-dimensional blade shape by radially stacking multiple two-dimensional blade profiles, which enables the blade to meet aerodynamic efficiency and the throat area to meet design targets.

[0058] In some embodiments, the step of measuring the actual thickness distribution of the coating in step 130 includes:

[0059] Take at least three sections along the blade height direction and cover the entire blade height, for example, it can still cover the inner diameter A of the flow channel, the outer diameter C of the flow channel, and the middle diameter B of the flow channel;

[0060] Two guide blades are cut along each cross section to obtain a two-dimensional blade profile. The coating thickness distribution on the surface of the two-dimensional blade profile at each cross section is measured to obtain the actual coating spraying thickness distribution.

[0061] In this process, after measuring the actual coating thickness distribution of multiple cross sections, when designing the second theoretical blade model, the actual coating thickness distribution can be designed according to the corresponding actual coating thickness distribution at a certain radial height cross section, or the coating thickness of the blade at a specific position in the circumferential direction can be averaged with the actual coating thickness of multiple cross sections.

[0062] This embodiment, by measuring the actual coating thickness distribution of the blade at different radial height positions, can more accurately design a second theoretical blade model, which can be used as the basic structure for spraying multi-blade products, resulting in multi-blade products that are more in line with aerodynamic efficiency after spraying.

[0063] In some embodiments, in the actual thickness distribution of the coating, there is a second coating thickness T2 in the shielded area between two adjacent first blade substrates, and a first coating thickness T1 in the unshielded area, wherein the second coating thickness T2 is less than the first coating thickness T1.

[0064] This embodiment involves spraying a coating onto at least two first blade substrates. The coating thickness is uniform on the surface of the unshielded areas, making it easy to ensure the first coating thickness T1. However, in the shielded areas, the coating thickness is difficult to reach the first coating thickness T1 due to the obstruction caused by the blades. By making the coating thinner in the shielded areas and compensating for the reduction in coating thickness on the surface of the second blade substrate in the shielded areas, the blade shape can still meet the design objectives.

[0065] In some embodiments, such as Figure 1 and Figure 5 As shown, a two-dimensional leaf profile is formed in the cross-section of the second blade substrate perpendicular to the blade height direction. The two-dimensional leaf profile includes the blade back 3 and the blade base 4. The position on the blade back 3 corresponding to the throat 7 is the second point P2. The first point P1 and the third point P3 are selected on both sides of the second point P2 along the blade back 3. The first point P1 and the third point P3 are located at the edge of the shading area. The coating thickness of the second point P2 is zero, and the actual coating thickness gradually increases from the second point P2 to the first point P1 and from the second point P2 to the third point P3 to the first coating thickness T1.

[0066] In this arrangement, the leaf base 4 of the first leaf 1 is positioned opposite the back 3 of the second leaf 2, with the obstruction point located at the trailing edge of the first leaf 1 and the leading edge of the second leaf 1. The position on the back 3 corresponding to the throat 7 is designated as point P2, and the distance between point P1 and point P2 is equal to the distance between point P3 and point P2. For example... Figure 3 As shown, the back 3 of the second blade 2 is provided with a throat line 8. A first blocking edge line 9 is provided at intervals on the side of the throat line 8 near the leading edge, and a second blocking edge line 10 is provided at intervals on the side of the throat line 8 near the trailing edge. A blocking area is formed between the first blocking edge line 9 and the second blocking edge line 10.

[0067] At the location of throat 7, i.e., within the preset range on both sides of the second point P2, the coating thickness is zero. Even after long-term service of the aero-engine and prolonged exposure of the high-pressure turbine to high-temperature combustion gases, there is no issue of coating peeling off at throat 7. This ensures that the area of ​​throat 7 matches the design target value, eliminating the sensitivity of throat area to coating thickness and preventing it from affecting the airflow pressure in the main flow channel. This not only prevents the main flow channel pressure from falling below the design value, which would lead to increased cooling sealing gas flow and increased aerodynamic mixing losses, thus improving aerodynamic efficiency, but also prevents the main flow channel pressure from exceeding the design value, which would reduce the operational reliability and service life of the blades and rim disk. Therefore, this type of turbine can improve operational reliability and service life.

[0068] Furthermore, the actual coating thickness gradually increases from the second point P2 to the first point P1 and from the second point P2 to the third point P3 to the first coating thickness T1. By setting transition zones on both sides of the throat 7, the area of ​​the coating on the blade back 3 that has not been coated can gradually transition to the first coating thickness T1, reducing the stress caused by the coating thickness change and reducing the risk of coating peeling off under the action of high-temperature combustion gas.

[0069] In some embodiments, such as Figure 1 and Figure 6As shown, a two-dimensional leaf profile is formed in the cross section of the second blade substrate perpendicular to the blade height direction. The two-dimensional leaf profile includes the blade back 3 and the blade base 4. The two ends of the shading area of ​​the blade base 4 are the fourth point P4 and the fifth point P5, respectively. The fourth point P4 is located at the tail edge. The position on the blade base corresponding to the throat is located between the fourth point P4 and the fifth point P5. The coating thickness at the fourth point P4 is zero. The actual coating thickness gradually increases from the fourth point P4 to the fifth point P5 to the first coating thickness T1.

[0070] In this design, the leaf base 4 of the first blade 1 is positioned opposite the back 3 of the second blade 2, with the obstruction located at the trailing edge of the first blade 1 and the leading edge of the second blade 1. Since a trailing edge slit is typically provided at the trailing edge of the first blade 1, the coating thickness at point P4 is zero to prevent the coating from peeling off. Figure 4 As shown, the leaf base 4 of the first blade 1 is provided with a trailing edge line 12 and a third shading edge line 11, and a shading area is formed between the third shading edge line 11 and the trailing edge line 12.

[0071] In this embodiment, the coating thickness at point P4 is zero, which prevents the coating from peeling off in the trailing edge area of ​​the blade. Since the area between point P4 and point P5 is blocked, the coating thickness in this area is less than the first coating thickness T1. The coating thickness gradually increases from point P4 to point P5 to the first coating thickness T1, which allows the blocked area and the unblocked area to transition smoothly at point P5, reducing the stress caused by the coating thickness change. Under the action of high-temperature combustion gas, the risk of coating peeling off can be reduced.

[0072] For example, a smooth curve can be used in the transition zone of the coating; in this transition zone, spline curves or polynomials can be used to describe the thickness distribution, and the thickness distribution is continuous of first order or even second order or higher, so that the surface of the casting model is smooth; moreover, it can prevent local discontinuity in coating thickness from causing large thermal stress and cracks under actual working conditions, reduce the stress on the coating in this area, and make it less likely to fall off, which is conducive to maintaining the durability of the coating.

[0073] Secondly, this disclosure provides a turbine, which in some embodiments includes at least one guide vane group arranged along the axial direction. Each guide vane group includes multiple blades arranged circumferentially, and at least two adjacent blades along the circumferential direction are connected to form a multi-stage blade. The multi-stage blade is formed according to the turbine multi-stage guide vane throat area control method of the above embodiments.

[0074] In some embodiments, such as Figure 7As shown, the blade has a rim plate 11. In a multi-blade configuration, the rim plates 11 of two adjacent blades are connected to each other without gaps at the connection interface (see position G1). In another configuration, the rim plates 11 of two adjacent blades are connected to each other with gaps at the connection interface (see position G2). For example, the rim plate 11 may be located on the side of the blade near the inner diameter A of the flow channel.

[0075] For example, the first-stage guide vane of a high-pressure turbine consists of dozens of guide vanes. Gas flow channels are formed between adjacent guide vanes, with the throat being the location of the minimum flow area and the throat area being the minimum flow area. Considering the thermal expansion and deformation that occurs during engine operation, gaps are designed in the circumferential direction between the rim plates 11 of adjacent vanes. Therefore, cold air will leak into the main flow channel through these circumferential gaps in the rim plates 11. Reducing this leakage will improve engine performance.

[0076] By employing a multi-blade structure, where two or more blades are grouped together and cast as a single unit, or by first casting individual guide blades and their flanges 11, and then welding the flanges 11 together to form a multi-blade structure with a seamless design, cold air leakage from the circumferential gaps in the guide blade flanges 11 can be reduced. Finally, a heat-insulating coating is sprayed onto the surface of the guide blades and flanges on the main flow side of the multi-blade structure, and the components are assembled into a complete ring.

[0077] In some embodiments, the blade assembly is a guide vane assembly. Turbine guide vanes require reduced cold air leakage; this design approach better enables the airfoil to meet aerodynamic efficiency requirements and ensures the throat area meets design objectives.

[0078] Furthermore, this disclosure provides an aero-engine including the turbine of the above embodiment. This embodiment can ensure the turbine's flow capacity and aerodynamic performance, thereby improving the performance and lifespan of the aero-engine.

[0079] The above are merely exemplary embodiments of this disclosure and are not intended to limit this disclosure. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of this disclosure should be included within the protection scope of this disclosure.

Claims

1. A method for controlling the throat area of ​​a multi-stage turbine guide vane, characterized in that, The turbine includes at least one set of guide vanes arranged along the axial direction, each set of guide vanes including a plurality of vanes arranged circumferentially, and at least two adjacent vanes along the circumferential direction are connected together to form a multi-blade assembly before spraying. Methods for controlling laryngeal area include: Based on the design blade shape that meets aerodynamic performance requirements, the shape is offset inward according to the uniform first coating thickness to form the first theoretical blade model. After the first blade substrate is formed according to the first theoretical blade model, at least two first blade substrates are connected into one piece, and there is a shielding area between two adjacent first blade substrates, and the throat of the blade is located in the shielding area. The coating is sprayed onto at least two of the first blade substrates according to the first coating thickness, and the actual thickness distribution of the coating is measured; Using the actual thickness distribution of the coating as the input for bias, based on the design blade shape that meets aerodynamic performance requirements, the shape is biased inward according to the actual thickness distribution to form a second theoretical blade model. After forming the second blade substrate according to the second theoretical blade model, at least two second blade substrates are connected together as one unit; The first coating thickness is applied to at least two second blade substrates to obtain the multi-blade product.

2. The method for controlling the throat area of ​​multi-stage turbine guide vanes according to claim 1, characterized in that, The steps involved in designing the shape of blades that meet aerodynamic performance requirements include: Select at least three radial height positions of the flow channel, covering at least the inner diameter, outer diameter, and middle diameter of the flow channel, and design two-dimensional airfoil profiles that meet aerodynamic performance, including the back of the blade and the blade base. The minimum distance between the back of the blade and the blade base of two adjacent two-dimensional airfoil profiles is the throat. By adopting the radial stacking law, the two-dimensional blade profiles with different flow channel radial heights are radially stacked to form a three-dimensional blade shape that satisfies aerodynamic performance.

3. The method for controlling the throat area of ​​multi-stage turbine guide vanes according to claim 1, characterized in that, The steps for measuring the actual thickness distribution of the coating include: Take at least three cross sections along the blade height direction, covering the entire blade height; Two guide blades are cut along each cross section to obtain a two-dimensional blade profile. The coating thickness distribution on the surface of the two-dimensional blade profile at each cross section is measured to obtain the actual coating spraying thickness distribution.

4. The method for controlling the throat area of ​​multi-stage turbine guide vanes according to claim 1, characterized in that, In the actual thickness distribution of the coating, there is a second coating thickness between two adjacent first blade substrates in the shielded area and the first coating thickness in the unshielded area, wherein the second coating thickness is less than the first coating thickness.

5. The method for controlling the throat area of ​​multi-stage turbine guide vanes according to any one of claims 1 to 4, characterized in that, A two-dimensional leaf profile is formed in the cross-section of the second leaf body perpendicular to the height direction of the leaf. The two-dimensional leaf profile includes the leaf back and the leaf base. The position on the leaf back corresponding to the throat is the second point (P2). A first point (P1) and a third point (P3) are selected on both sides of the second point (P2) along the leaf back. The first point (P1) and the third point (P3) are located at the edge of the shading area. The coating thickness at the second point (P2) is zero, and the actual coating thickness gradually increases from the second point (P2) to the first point (P1) and from the second point (P2) to the third point (P3) to the first coating thickness.

6. The method for controlling the throat area of ​​multi-stage turbine guide vanes according to any one of claims 1 to 4, characterized in that, A two-dimensional leaf profile is formed in the cross-section of the second leaf body perpendicular to the leaf height direction. The two-dimensional leaf profile includes the leaf back and the leaf base. The two ends of the shading area of ​​the leaf base are the fourth point (P4) and the fifth point (P5), respectively. The fourth point (P4) is located at the tail edge. The position on the leaf base corresponding to the throat is located between the fourth point (P4) and the fifth point (P5). The coating thickness at the fourth point (P4) is zero. The actual coating thickness gradually increases from the fourth point (P4) to the fifth point (P5) to the first coating thickness.

7. A turbine, characterized in that, The turbine multi-stage guide vane includes at least one set of guide vanes arranged along the axial direction, each set of guide vanes includes multiple vanes arranged circumferentially, and at least two adjacent vanes in the circumferential direction are connected to form a multi-stage vane, wherein the multi-stage vane is formed in accordance with the method for controlling the throat area of ​​turbine multi-stage guide vanes according to any one of claims 1 to 6.

8. The turbine according to claim 7, characterized in that, The blade has a rim plate (11), and the rim plates (11) of two adjacent blades in the multi-section blade are connected to each other and there is no gap at the connection interface. The rim plates (11) of two adjacent blades in the multi-section blade are connected to each other and there is a gap at the connection interface.

9. The turbine according to claim 7, characterized in that, The blade assembly is a guide blade assembly.

10. An aircraft engine, characterized in that, Includes the turbine as described in any one of claims 7 to 9.

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