A design method for CMC turbine rotor blade simulation parts

Abstract

The present invention discloses a CMC turbine rotor blade airfoil simulation design method, comprising the following steps: the first step is to establish a complete CMC turbine rotor blade macro-geometry model; the second step is to perform finite element analysis of the complete blade under given working conditions to obtain stress distribution; the third step is to extract the blade profile of the dangerous area of ​​the blade airfoil as the blade profile of the simulation test section; the fourth step is to establish a blade airfoil simulation geometry model; the fifth step is to use a progressive damage algorithm to obtain the stress distribution of the blade airfoil simulation; the sixth step is to evaluate the consistency of the stress distribution of the simulation and the complete blade dangerous area. The present invention makes up for the current situation of the lack of CMC rotor blade airfoil simulation design method, provides an effective means for the assessment of the process and strength scheme of the blade, and helps to promote the engineering application of CMC in turbine rotor blades.

Description

Technical Field

[0001] The invention belongs to the technical field of aircraft engine turbine blades, and in particular relates to a CMC turbine rotor blade blade body simulation component design method. Background Art

[0002] Ceramic matrix composites (CMC) have low density (about 1 / 3 of traditional high-temperature alloys), high temperature resistance and oxidation resistance. The application of ceramic matrix composites can increase the temperature of the hot end components of the engine to 1200℃~1350℃, making it an ideal high-temperature material in the aerospace field.

[0003] As one of the most important hot-end components, the turbine rotor blades must not only consider their static strength but also their fatigue failure during the design process. The loads on the actual turbine rotor blades during operation are very complex. They are not only subjected to centrifugal tensile loads, but also cause high-frequency vibrations due to airflow excitation during flight. For the more complex structure of CMC turbine rotor blades, each characteristic interval varies greatly due to the manufacturing process and the loads they are subjected to. If tests are carried out simultaneously in the early stage of the design, it will be impossible to quantify some influencing factors due to too many coupling factors.

[0004] Therefore, it is an effective method to design CMC turbine rotor blades by designing reasonable characteristic structure simulation parts for different characteristic parts of the blade and studying the failure mode of characteristic structure simulation parts under specific loads. In recent years, many scholars have evaluated the failure mode of the tenon part of CMC turbine rotor blades by designing characteristic simulation parts. The damage evolution process can be studied by conducting static tensile tests and tension-tension fatigue tests on CMC tenon simulation parts. However, for the blade body part of turbine rotor blades, which also has a high risk of fatigue failure, few scholars have carried out simulation part design for this characteristic area, and it is urgent to study its damage evolution process and failure mode.

[0005] In summary, there is currently a lack of simulation parts design methods for the blade body part of CMC turbine rotor blades, and it is necessary to conduct research on this. Summary of the invention

[0006] The technical problem to be solved by the present invention is to meet the evaluation needs of the static strength and fatigue failure mode of the blade body of a CMC turbine rotor blade, and to provide a blade body simulation part design method that is consistent with the actual CMC blade profile process and can reflect the stress distribution characteristics of the dangerous cross-section of the CMC turbine rotor blade blade body, thereby making up for the current lack of a CMC rotor blade blade body simulation part design method.

[0007] In order to achieve the above technical objectives, the technical solution adopted by the present invention is:

[0008] A CMC turbine rotor blade airfoil simulation design method comprises the following steps:

[0009] Step S1: establishing a macro-geometric model of a CMC turbine rotor blade;

[0010] Step S2: assigning macro-equivalent anisotropic thermal conductivity parameters and mechanical performance parameters to the macro-geometric model, performing finite element analysis of the structural performance of the CMC turbine rotor blade under given working conditions, and obtaining the stress distribution of the macro-geometric model;

[0011] Step S3: extracting the blade profile of the dangerous area of ​​the blade body in the macro-geometric model of the turbine rotor blade based on the stress distribution of the macro-geometric model;

[0012] Step S4: establishing a blade airfoil simulation geometric model, wherein the blade airfoil simulation geometric model comprises a test section, a clamping section and a transition section, wherein both ends of the test section along the axial direction of the tensile load are connected to the clamping section through the transition section, respectively, and the blade shape of the blade airfoil dangerous area is the blade shape of the test section;

[0013] Step S5: assigning anisotropic mechanical performance parameters to the blade airfoil simulation geometric model, obtaining the stress distribution of the blade airfoil simulation geometric model under a predetermined tensile load based on a progressive damage algorithm, and performing a high stress area check on the blade airfoil simulation geometric model to avoid the high stress area appearing in the transition section;

[0014] Step S6: Compare the cross-sectional stress distribution of the test section of the blade simulation part with the cross-sectional stress distribution of the blade dangerous area in the macro-geometric model of the turbine rotor blade, and adjust the relative position of the test section of the blade simulation part and the clamping section to make the cross-sectional stress distribution of the test section of the blade simulation part consistent with the cross-sectional stress distribution of the blade dangerous area in the macro-geometric model of the turbine rotor blade, thereby completing the design of the blade simulation part.

[0015] To optimize the technical solution, the present invention is further improved as follows:

[0016] In step S2, the macroscopic equivalent anisotropic thermal conductivity parameters and mechanical performance parameters include thermal conductivity, material density, thermal expansion coefficient, elastic modulus, Poisson's ratio and material tensile failure strength.

[0017] In step S2, the given working condition is to simulate the actual service condition of the CMC turbine rotor blade, which is subjected to three loads of inertial load, aerodynamic load and thermal stress simultaneously during the service process.

[0018] The stress distribution of the macro-geometric model is the normal stress distribution along the blade height direction on the CMC turbine rotor blade.

[0019] In step S3, the blade airfoil dangerous area refers to the area with the largest stress in the stress distribution of the blade airfoil part of the CMC turbine rotor blade.

[0020] The clamping section is the clamping area of ​​the fixture during the geometric model test of the blade simulation part, and the transition section connects the test section and the clamping section in a smooth transition manner.

[0021] One side of the cross section of the test section perpendicular to the axial direction of the tensile load corresponds to the leading edge of the blade, and the other side corresponds to the trailing edge of the blade. One side of the clamping section is provided with an arc-shaped edge, and the arc-shaped edge is on the same side as the trailing edge of the blade.

[0022] The method for establishing the geometric model of the blade simulation part is specifically as follows: first, the two-dimensional test section cross-section blade shape is stretched along the axial direction of the tensile load to obtain a three-dimensional test section; then a two-dimensional clamping section cross-section curve is established, and the clamping section cross-section curve is stretched along the axial direction of the tensile load to obtain a three-dimensional clamping section; then, the same number of isoparametric curves are established on the test section and the side of the clamping section, and the isoparametric curves at corresponding positions are connected using bridging curves, and finally, the geometric model of the blade simulation part is obtained by stitching all the bridging curves through an artistic surface, wherein the artistic surface after stitching all the bridging curves is a transition section.

[0023] In step S5, the anisotropic mechanical performance parameters assigned to the geometric model of the blade simulation component include material density, thermal expansion coefficient, elastic modulus, Poisson's ratio, and material tensile failure strength. The specific values ​​of the anisotropic mechanical performance parameters are consistent with the specific values ​​of the macroscopic equivalent anisotropic thermal conductivity parameters and mechanical performance parameters in step S2.

[0024] The beneficial effect of the present invention is that the traditional strength and life assessment of CMC turbine rotor blades usually focuses on the tenon area with a higher risk of failure. However, in the actual service of CMC turbine rotor blades, the blade body part also has a high risk of failure due to the coupling of multiple loads, and the weaving process and stress distribution of the blade body part are relatively complex, which brings great difficulty to the test assessment. The present invention extracts the geometry and process characteristics of the dangerous section of the blade body of a complete CMC turbine rotor blade, simplifies the complete blade, and proposes a CMC blade body simulation part design method that is consistent with the process of the dangerous section of the blade body of a real blade, and can reflect the stress gradient distribution characteristics of the dangerous section of the blade body under simple axial loading conditions. The CMC blade body simulation part can be used to implement the strength life test of the dangerous section of the blade body of a CMC turbine rotor blade, providing an effective means for the process and strength scheme assessment of the blade, which is helpful to promote the engineering application of CMC in turbine rotor blades. BRIEF DESCRIPTION OF THE DRAWINGS

[0025] Figure 1 It is a schematic diagram of the design process of CMC turbine rotor blade airfoil simulation parts;

[0026] Figure 2 is a stress distribution diagram of a complete blade along the warp direction in the embodiment;

[0027] Figure 3 is a stress distribution diagram of a dangerous section of a blade body of a CMCs low-pressure turbine rotor blade in an embodiment;

[0028] Figure 4 Schematic diagram of the relative position relationship between the blade body simulation test section and the clamping section in the embodiment;

[0029] Figure 5 : is a stress distribution diagram of the blade body simulation test section in the embodiment;

[0030] Figure 6 is a dimension diagram of a blade body simulation component established in the embodiment;

[0031] Figure 7 It is a schematic diagram of the positions of the isoparametric curve and the bridging curve established in the embodiment. DETAILED DESCRIPTION

[0032] In order to make the purpose, technical solutions and advantages of the present application clearer, the present application is described and illustrated below in conjunction with the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are only used to explain the present application and are not intended to limit the present application. Based on the embodiments provided in the present application, all other embodiments obtained by ordinary technicians in the field without making creative work are within the scope of protection of the present application.

[0033] Obviously, the drawings described below are only some examples or embodiments of the present application. For ordinary technicians in this field, the present application can also be applied to other similar scenarios based on these drawings without creative work. In addition, it can also be understood that although the efforts made in this development process may be complicated and lengthy, for ordinary technicians in this field related to the content disclosed in this application, some changes in design, manufacturing or production based on the technical content disclosed in this application are just conventional technical means, and should not be understood as insufficient content disclosed in this application.

[0034] Reference to "embodiments" in this application means that a particular feature, structure, or characteristic described in conjunction with the embodiments may be included in at least one embodiment of the present application. The appearance of the phrase in various locations in the specification does not necessarily refer to the same embodiment, nor is it an independent or alternative embodiment that is mutually exclusive with other embodiments. It is explicitly and implicitly understood by those of ordinary skill in the art that the embodiments described in this application may be combined with other embodiments without conflict.

[0035] Unless otherwise defined, the technical terms or scientific terms involved in this application should be understood by people with ordinary skills in the technical field to which this application belongs. The words "one", "a", "a", "the" and the like involved in this application do not indicate a quantitative limitation, and may represent the singular or plural. The terms "include", "comprise", "have" and any of their variations involved in this application are intended to cover non-exclusive inclusions; for example, a process, method, system, product or device that includes a series of steps or units (units) is not limited to the listed steps or units, but may also include steps or units that are not listed, or may also include other steps or units inherent to these processes, methods, products or devices. The words "connect", "connected", "coupled" and the like involved in this application are not limited to physical or mechanical connections, but may include electrical connections, whether direct or indirect. The "multiple" / "several" involved in this application refers to two or more. "And / or" describes the association relationship of associated objects, indicating that there can be three relationships, for example, "A and / or B" can mean: A exists alone, A and B exist at the same time, and B exists alone. The character " / " generally indicates that the objects before and after are in an "or" relationship. The terms "first", "second", "third", etc. involved in this application are only used to distinguish similar objects and do not represent a specific ordering of the objects.

[0036] like Figure 1 As shown, the present invention discloses a design and test method for a woven ceramic matrix composite (CMC) turbine rotor blade airframe simulation component, comprising the following steps:

[0037] The first step is to establish a complete CMC turbine rotor blade geometry model (including blade crown, edge plate, tenon and other characteristic structures), and layer the model according to the actual prefabricated body direction.

[0038] In the second step, the established CMC turbine rotor blade geometric model is imported into the finite element analysis software, the macro-equivalent anisotropic thermal conductivity parameters and mechanical performance parameters of the CMC turbine rotor blade are given, and the unit local coordinate system of the model is defined; the CMC turbine rotor blade is subjected to thermal stress, aerodynamic bending moment and inertial load at the same time, and the flow heat transfer calculation is first performed in the flow field analysis software, and then the overall temperature field of the blade and the pressure field on the blade surface are used as thermal loads and aerodynamic loads to perform finite element analysis on the structural performance of the CMC turbine rotor blade, and the following is obtained: Figure 2 The stress distribution along the warp direction of the CMC turbine rotor blade under actual service conditions is shown in Figure 3 The figure shows the stress distribution diagram of the dangerous section of the blade of the CMCs low-pressure turbine rotor blade. It can be seen from the figure that the stress at the dangerous section gradually decreases from the blade basin to the blade back, and gradually changes from tensile load to compressive load, presenting a gradient distribution.

[0039] In the third step, based on the calculation results of the second step, the blade shape of the dangerous area of ​​the CMC turbine rotor blade body is selected as Figure 6 The blade profile of the blade simulation test section shown in the AA section.

[0040] The fourth step is to establish a blade simulation geometric model based on the blade profile of the dangerous section extracted in the third step. First, the blade profile of the test section is stretched to obtain the following Figure 6 The test section shown is of length a mm. Figure 6 In the figure, the length of the transition section is b mm, the length of the clamping section is c mm, d is the maximum inscribed circle diameter of the blade profile of the test section section, f is the width of the clamping section, e is the thickness of the clamping section, and r is the radius of the arc-shaped edge.

[0041] In the design process of CMC turbine rotor blades, in order to ensure a smooth transition between the root and the tenon, the center point of the chord of the blade root blade profile is usually made to coincide with the midpoint of the geometric center line of the upper surface of the tenon. However, in the design process of the blade body simulation part, it was found that since the load on the simulation part is an axial tensile load, if the design is carried out according to the above design principles, the stress distribution of the simulation part test section will be quite different from the stress distribution of the dangerous section of the blade. The maximum stress point of the simulation part test section often appears on the back of the simulation part and the trailing edge point. In order to make the stress distribution of the cross-section of the simulation part test section similar to the dangerous section of the complete blade body, the present invention proposes a blade body simulation part design scheme based on eccentric design, such as Figure 4 The figure shows the relative positions of the clamping section section of the simulation component and the test section section (the dotted line in the figure is the geometric center line of the section, point E and point F are the front (rear) edge points of the test section blade, and point C and point D are the two end points on the clamping section section. During the design process, point E is made close to point C, and point F is made close to point D), so that the geometric center of the test section section of the simulation component deviates from the geometric center line of the clamping section section. At the same time, the right side of the clamping section section is designed to be a smooth arc to weaken the stress concentration at the trailing edge of the blade of the test section of the simulation component. Finally, by continuously fine-tuning the relative positions of the two, the stress distribution of the test section of the simulation component can meet the requirements.

[0042] Based on the above principle of relative position of cross sections, the cross-section curve of the clamping section is established and stretched to obtain the clamping section. Then, the same number of isoparametric curves are established on the sides of the test section and the clamping section model, and the isoparametric curves at the corresponding positions are connected using bridge curves. Finally, all bridge curves are stitched together through the art surface to obtain the transition section model that smoothly connects the test section and the clamping section.

[0043] The fifth step is to import the established CMC blade simulation component geometric model into the finite element analysis software, assign it macro-equivalent anisotropic thermal conductivity parameters and mechanical performance parameters (the specific values ​​of the parameters are consistent with the parameters in the second step), and define the local coordinate system of the model unit; in the finite element software, apply a fixed support constraint to one end of the simulation model and an axial tensile load to the other end to perform finite element analysis of the structural performance of the blade simulation component.

[0044] like Figure 3 The figure shows the stress distribution diagram of the dangerous section of the blade of the CMCs low-pressure turbine rotor. Figure 5 The stress distribution diagram of the test section of the blade simulation designed for the present invention shows that the stress distribution forms of the two are exactly the same, the maximum stress points appear at the leading and trailing edges of the blade, and the stress level gradually decreases from the blade basin to the blade back and finally turns into a compression load. In addition to the similar stress distribution forms, the stress ranges of the two are also basically the same, both of which are about -50MPa to 180MPa. It can be seen that the blade simulation designed for the present invention has good similarity.

[0045] The above are only preferred embodiments of the present invention, and the protection scope of the present invention is not limited to the above embodiments. All technical solutions under the concept of the present invention belong to the protection scope of the present invention. It should be pointed out that for ordinary technicians in this technical field, some improvements and modifications without departing from the principle of the present invention should be regarded as the protection scope of the present invention.

Claims

1. A CMC turbine rotor blade airfoil simulation design method, characterized in that: The following steps are involved: Step S1: establishing a macro-geometric model of a CMC turbine rotor blade; Step S2: assigning macro-equivalent anisotropic thermal conductivity parameters and mechanical performance parameters to the macro-geometric model, performing finite element analysis of the structural performance of the CMC turbine rotor blade under given working conditions, and obtaining the stress distribution of the macro-geometric model; Step S3: extracting the blade profile of the dangerous area of ​​the blade body in the macro-geometric model of the turbine rotor blade based on the stress distribution of the macro-geometric model; Step S4: establishing a blade airfoil simulation geometric model, wherein the blade airfoil simulation geometric model comprises a test section, a clamping section and a transition section, wherein both ends of the test section along the axial direction of the tensile load are connected to the clamping section through the transition section, respectively, and the blade shape of the blade airfoil dangerous area is the blade shape of the test section; Step S5: assigning anisotropic mechanical performance parameters to the blade airfoil simulation geometric model, obtaining the stress distribution of the blade airfoil simulation geometric model under a predetermined tensile load based on a progressive damage algorithm, and performing a high stress area check on the blade airfoil simulation geometric model to avoid the high stress area appearing in the transition section; Step S6: comparing the stress distribution of the cross section of the test section of the blade airfoil simulation part with the stress distribution of the cross section of the blade airfoil danger area in the macro-geometric model of the turbine rotor blade, and adjusting the relative position of the test section of the blade airfoil simulation part and the clamping section so that the stress distribution of the cross section of the test section of the blade airfoil simulation part is consistent with the stress distribution of the cross section of the blade airfoil danger area in the macro-geometric model of the turbine rotor blade, thereby completing the design of the blade airfoil simulation part; The method for establishing the geometric model of the blade simulation part is specifically as follows: first, the two-dimensional test section cross-section blade shape is stretched along the axial direction of the tensile load to obtain a three-dimensional test section; then a two-dimensional clamping section cross-section curve is established, and the clamping section cross-section curve is stretched along the axial direction of the tensile load to obtain a three-dimensional clamping section; then, the same number of isoparametric curves are established on the test section and the side of the clamping section, and the isoparametric curves at corresponding positions are connected using bridging curves, and finally, the geometric model of the blade simulation part is obtained by stitching all the bridging curves through an artistic surface, wherein the artistic surface after stitching all the bridging curves is a transition section.

2. A CMC turbine rotor blade airfoil simulation design method according to claim 1, characterized in that: In step S2, the macroscopic equivalent anisotropic thermal conductivity parameters and mechanical performance parameters include thermal conductivity, material density, thermal expansion coefficient, elastic modulus, Poisson's ratio and material tensile failure strength.

3. A CMC turbine rotor blade airfoil simulation design method according to claim 1, characterized in that: In step S2, the given working condition is to simulate the actual service condition of the CMC turbine rotor blade, which is subjected to three loads of inertial load, aerodynamic load and thermal stress simultaneously during the service process.

4. A CMC turbine rotor blade airfoil simulation design method according to claim 1, characterized in that: The stress distribution of the macro-geometric model is the normal stress distribution along the blade height direction on the CMC turbine rotor blade.

5. A CMC turbine rotor blade airfoil simulation design method according to claim 1, characterized in that: In step S3, the blade airfoil dangerous area refers to the area with the largest stress in the stress distribution of the blade airfoil part of the CMC turbine rotor blade.

6. A CMC turbine rotor blade airfoil simulation design method according to claim 1, characterized in that: The clamping section is the clamping area of ​​the fixture during the geometric model test of the blade simulation part, and the transition section connects the test section and the clamping section in a smooth transition manner.

7. A CMC turbine rotor blade airfoil simulation design method according to claim 1, characterized in that: One side of the cross section of the test section perpendicular to the axial direction of the tensile load corresponds to the leading edge of the blade, and the other side corresponds to the trailing edge of the blade. One side of the clamping section is provided with an arc-shaped edge, and the arc-shaped edge is on the same side as the trailing edge of the blade.

8. A CMC turbine rotor blade airfoil simulation design method according to claim 1, characterized in that: In the step S5, the anisotropic mechanical performance parameters assigned to the geometric model of the blade simulation component include material density, thermal expansion coefficient, elastic modulus, Poisson's ratio, and material tensile failure strength. The specific values ​​of the anisotropic mechanical performance parameters are consistent with the specific values ​​of the macroscopic equivalent anisotropic thermal conductivity parameters and mechanical performance parameters in step S2.

Images