Integrated blade root skin fatigue life analysis method

Abstract

The invention belongs to the technical field of rotor strength design, and particularly relates to an integrated blade root skin fatigue life analysis method. Comprising the following steps: step 1, determining a fatigue dangerous area by analyzing a blade root load transmission path; 2, extracting a simplified blade root model according to the fatigue dangerous area determined in the step 1; 3, constructing a blade root finite element model according to the blade root skin model in the step 2; 4, calculating strain distribution of the finite element model of the blade root by using a finite element method, and further identifying a strain dangerous point; and 5, determining a static load working condition and a dynamic load working condition by combining the flight spectrum and the load spectrum, and analyzing the service life of the root section according to a safe service life analysis method by utilizing the static strain of the static load working condition and the dynamic strain of the dynamic load working condition obtained by a finite element method. The defects of a traditional engineering method in the aspects of blade root tapping position and skin strain calculation are overcome.

Description

Technical Field

[0001] This invention belongs to the field of rotor strength design technology, specifically relating to a method for analyzing the fatigue life of an integrated blade root skin. Background Technology

[0002] Rotor blades play a crucial role in helicopters and vertical takeoff and landing aircraft, and their weight and cost affect the safety and economy of the entire aircraft.

[0003] In recent years, the rise of vertical takeoff and landing aircraft, represented by eVTOL, has brought an integrated blade structure into the view of researchers.

[0004] This integrated rotor consists of two blades, left and right, with a 180-degree phase difference. It is a hubless, two-bladed, rigid, fixed-pitch rotor. Through openings in the blade root skin, connecting bolts pass through the blade skin, fiber packing, metal cover, and other related components before directly connecting to the motor shaft. Compared to the complex structure and power transmission path of blades and hubs in traditional rotor systems, this configuration is simpler, has fewer parts, lower cost, and a more transparent power transmission path.

[0005] Currently, most helicopter rotor blades still use a wound spars structure at the root. However, integrated rotor blades achieve structural connection through openings in the root skin. Therefore, the traditional roving-wound rotor blade root strength calculation method based on materials mechanics is no longer applicable in order to calculate the strain field of the root skin under load and the stress concentration at the edge of the opening. Therefore, there is an urgent need to develop a method suitable for fatigue life analysis of the root skin of integrated rotor blades. Summary of the Invention

[0006] Purpose of the invention: This invention provides a fatigue life analysis method for the root skin of an integrated rotor blade, which effectively obtains the location of the opening at the root of the integrated rotor blade and the strain field of the skin after loading. It makes up for the shortcomings of traditional engineering methods in calculating the location of the opening at the root of the rotor blade and the strain of the skin, and provides a reference for the design of other helicopter rotor systems that adopt integrated rotor blade configurations.

[0007] To address the aforementioned technical issues, a fatigue life analysis method for the root skin of an integrated blade is provided. This method is applied to an integrated blade, which includes a first blade and a second blade that are centrally symmetrical and 180 degrees out of phase. The blade root is formed at the junction of the first and second blades. The blade root is filled with fiber filler blocks. The blade root has through holes that penetrate the skin and fiber filler blocks. The upper and lower skin surfaces of the blade root are respectively covered with an upper metal cover plate and a lower metal cover plate. Using connecting bolts, the upper metal cover plate, blade skin, fiber filler blocks, and lower metal cover plate are sequentially connected to the motor shaft.

[0008] The fatigue life analysis method includes the following steps: Step 1: Identify fatigue-prone areas by analyzing the load transfer path at the blade root; Step 2: Extract a simplified blade root model based on the fatigue-prone area identified in Step 1; Step 3: Construct a finite element model of the blade root based on the blade root skin model in Step 2; Step 4: Calculate the strain distribution of the finite element model at the blade root using the finite element method to further identify strain hazard points; Step 5: Combining the flight spectrum and load spectrum, determine the static load condition and dynamic load condition. Using the static strain of the static load condition and the dynamic strain of the dynamic load condition obtained by the finite element method, analyze the life of the root section according to the safe life analysis method.

[0009] In one possible embodiment, the specific process of load transfer path analysis and fatigue hazard area determination in step one is as follows: Centrifugal force is transmitted to the blade root through the blade skin, then to the motor shaft through the connecting bolts, and finally balances at the center of the motor shaft. Waving moment is transmitted to the blade root and metal cover plate through the blade skin, then to the motor shaft through the connecting bolts. The static waving moment can be balanced at the center of the motor shaft, but the dynamic waving moment is superimposed at the center of the motor shaft. Shivering moment is transmitted to the blade root through the blade skin, then to the metal cover plate through the friction between the blade root skin and the metal cover plate, and finally to the motor shaft through the connecting bolts. The shivering moment is balanced at the center of the motor shaft, but the static shivering moment is superimposed at the center of the motor shaft. Based on the above load transmission path analysis, the fatigue hazard area is located in the contact area between the metal cover plate and the skin, the transition area of ​​the skin, and the opening area of ​​the blade root skin. The transition area of ​​the skin refers to the area extending from the blade root towards the tips of the first and second blades.

[0010] In one possible embodiment, the specific process of extracting the simplified blade root model in step two includes: Extract the model including the contact area between the metal cover plate and the skin, the skin opening area at the blade root, and the transition area of ​​the skin; The blade skin is made of multiple layers of glass cloth or carbon cloth, and the outer surface of the blade skin contour is extracted to replace the complex layered structure. The connecting bolts, upper metal cover plate, and lower metal cover plate that mate with the blade root skin are extracted directly without geometric simplification, maintaining their assembly relationship.

[0011] In one possible embodiment, step three specifically includes the following steps: Material parameters and constitutive definitions are defined for the blade root skin, upper metal cover plate, lower metal cover plate, and fiber filler block in the simplified blade root model. The simplified blade root model is meshed to obtain the finite element model of the blade root. Set the load and displacement boundary conditions; Configure the interaction relationships of each component in the finite element model of the blade root.

[0012] In one possible embodiment, in step four, the strain calculation of the finite element model at the blade root is simplified to a plane stress problem, and the specific calculation is performed according to the following steps: Calculate the element stiffness matrix and the element equivalent nodal load array; Calculate the structural stiffness matrix and the structural nodal load array; Calculate the structural displacement solution; Calculate the structural strain.

[0013] In one possible embodiment, the specific process of identifying strain hazard points based on strain distribution in step four includes: The maximum strain point and strain red zone are obtained by statistical analysis of the structural strain field, which are used as the identified strain hazard points.

[0014] In one possible embodiment, in step five, the safe life analysis includes high-cycle fatigue life analysis, low-cycle fatigue life analysis, and total fatigue life analysis.

[0015] In one possible embodiment, the specific process of high-cycle fatigue life analysis includes: Calculate the equivalent strain of high-cycle fatigue; Calculate the number of high-frequency safety cycles; Calculate high-cycle damage.

[0016] In one possible embodiment, the specific process of low-cycle fatigue life analysis includes: Calculate low-cycle fatigue strain; Calculate the number of high-frequency safety cycles; Calculate low-cycle damage.

[0017] In one possible embodiment, the specific process of fatigue total life analysis includes: Calculate the total damage (high-cycle damage + low-cycle damage). Calculate the total lifetime.

[0018] In summary, the beneficial effects of the present invention are as follows: This invention proposes a finite element method for fatigue life analysis of integrated blade root skin. Compared with the traditional engineering method for calculating the strength of blade root using coarse sand winding based on mechanics of materials, this method considers the load redistribution caused by blade root deformation, provides a solution for statically indeterminate problems with multiple bolts fixing the blade root, establishes a high-precision finite element model of the integrated blade root skin, effectively obtains the strain field of the integrated blade root skin under load, accurately analyzes the stress concentration in the opening area of ​​the blade root skin and the complex surface load transfer problem between the cover plate and the skin, and makes up for the shortcomings of mechanics of materials methods in studying stress concentration at the opening location of the blade root skin, skin strain calculation, and surface load transfer. It provides a reference for the design of other helicopter rotor systems using integrated blade root skin opening configurations. Attached Figure Description

[0019] Figure 1 This is a flowchart of a preferred embodiment of the present invention; Figure 2A This is a schematic diagram of the integrated blade structure of a preferred embodiment of the present invention; Figure 2B This is a cross-sectional view of the integrated blade structure of a preferred embodiment of the present invention; Among them, ①: connecting bolts ②: cover plate ③: foam ④: blade skin ⑤: chopped fiber packing Figure 3 This is a schematic diagram of the integrated blade root model of a preferred embodiment of the present invention; Figure 4 This is a schematic diagram of the integrated blade root skin mesh of a preferred embodiment of the present invention; Figure 5 This is a schematic diagram of the cover plate mesh in the finite element model of the blade root of the preferred embodiment of the present invention; Figure 6 This is a schematic diagram of the chopped fiber packing block in the finite element model of the blade root of a preferred embodiment of the present invention; Figure 7 This is a schematic diagram of the foam mesh in the finite element model of the blade root of a preferred embodiment of the present invention; Figure 8 This is a schematic diagram of the integrated blade root loading of a preferred embodiment of the present invention (F). c Centrifugal force, M b : Swinging moment, M t (Swing moment); Figure 9 Domain Ω, subdomain Ωe, boundary S σ Boundaries of subdomains Schematic diagram; Figure 10 Meridional strain under static load on integrated blade root skin εs ; Figure 11 Meridional strain of integrated blade root skin under dynamic load ε d . Detailed Implementation

[0020] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0021] The features and illustrative embodiments of various aspects of the present invention will now be described in detail. Numerous specific details are set forth in the following detailed description to provide a thorough understanding of the invention. However, it will be apparent to those skilled in the art that the invention may be practiced without requiring some of these specific details. The following description of embodiments is merely intended to provide a better understanding of the invention by illustrating examples of the invention. The invention is by no means limited to any specific setups and methods set forth below, but covers any improvements, substitutions, and modifications to structures, methods, and devices without departing from the spirit of the invention. Well-known structures and techniques are not shown in the drawings and the following description to avoid unnecessarily obscuring the invention.

[0022] It should be noted that, unless otherwise specified, the embodiments of the present invention and the features thereof can be combined with each other, and the various embodiments can be referenced and cited in each other. The present invention will now be described in detail with reference to the accompanying drawings and embodiments.

[0023] The following example, using the strength design of the root section of a certain type of eVTOL integrated lift propeller blade, further illustrates the invention in detail. The method for analyzing the fatigue life of the integrated blade root skin comprises the following steps: 1. Analyze the load transfer path and locate the fatigue hazard area; The integrated blade of a certain eVTOL model primarily provides lift; the root consists of blade skin, connecting bolts, a metal cover plate, chopped fiber filler, and a bushing; the assembly method involves passing the connecting bolts through the blade skin, chopped fiber filler, and upper metal cover plate, and then fixing them to the motor shaft (see...). Figures 2A-2BCentrifugal force is transmitted through the blade skin to the blade root, then through the connecting bolts to the motor shaft, and finally balances at the center of the motor shaft. Waving moment is transmitted through the blade skin to the blade root and metal cover plate, then through the connecting bolts to the motor shaft. It's important to note that while the static waving moment can balance at the center of the motor shaft, the dynamic waving moment needs to be superimposed at the center. Shivering moment is transmitted through the blade skin to the blade root, then through the friction between the blade root skin and the metal cover plate to the metal cover plate, and finally through the connecting bolts to the motor shaft. It's important to note that while the shivering moment can balance at the center of the motor shaft, the static shivering moment needs to be superimposed at the center. Based on the above load transmission path analysis, the fatigue hazard areas can be roughly located in the contact area between the cover plate and the skin, the transition area of ​​the skin, and the opening area at the root of the skin.

[0024] 2. Extract and simplify the integrated blade root model The integrated blade root model includes the blade root skin and other components that cooperate with the blade root skin (as shown in Figure 2).

[0025] To accurately analyze the stress concentration and strain-prone areas at the edge of the skin holes at the root of the integrated blade, and to reduce subsequent calculations and improve computational efficiency, it is necessary to extract the digital models of the blade skin root section, transition section, and a small portion of the airfoil section as the research objects. Furthermore, since the blade skin is composed of multiple layers of glass cloth or carbon cloth, to simplify the model extraction process, the outer surface of the blade skin contour can be extracted instead of the complex layered structure.

[0026] Other components that mate with the blade root skin can be extracted directly without geometric simplification, maintaining their assembly relationships.

[0027] The simplified integrated blade root model is shown below. Figure 3 .

[0028] 3. Establish an integrated finite element model of the blade root. The following steps are required to establish a finite element model: a) geometric correction of the blade root model; b) material parameter setting; c) mesh generation; d) boundary condition setting; e) setting the interaction relationships between components.

[0029] a. Import the simplified blade root model into the finite element analysis software. Sometimes the model contains many complex surfaces and some modeling defects. During the import process, after data conversion, many tiny gaps, tiny steps, and some small geometric features will appear. At this time, it is necessary to use the geometric repair tool to correct the blade root model.

[0030] b. Define the material parameters and constitutive models for the blade root skin, cover plate, chopped fiber filler, and foam filling. The composite material of the blade root skin has poor plasticity; therefore, a linear elastic constitutive model is generally used to characterize the strain relationship. The strain analysis of the blade skin can be simplified to a plane stress laminate problem; therefore, only four independent elastic constants need to be given for the material parameters: E1 (meridian elastic modulus), E2 (zonal elastic modulus), ν12 (Poisson's ratio), and G12 (in-plane shear modulus). The metal cover plate generally uses isotropic steel or aluminum; only the elastic modulus and Poisson's ratio need to be given for the material parameters. The chopped fiber filler uses isotropic chopped fibers; only the elastic modulus and Poisson's ratio need to be given for the material parameters. The foam filling uses isotropic foam; only the elastic modulus and Poisson's ratio need to be given for the material parameters.

[0031] c. Mesh the blade root model. The blade skin is meshed using traditional quadrilateral shell elements, and the meshing result is as follows. Figure 4 As shown, the cover plate, chopped fiber filler, and foam were meshed using second-order tetrahedral elements. The meshing result is as follows. Figure 5 , Figure 6 and Figure 7 As shown; d. Set the load and displacement boundary conditions for the divided finite element model. First, set the load boundaries. The root of the integrated blade mainly bears the flapping bending moment and the centrifugal force under high-speed rotation. The flapping bending moment is divided into static and dynamic moments. The static bending moment remains unchanged in direction and magnitude within one rotation cycle, while the dynamic bending moment changes periodically in magnitude and direction according to a cosine function within one rotation cycle, which can be expressed by the formula: Swinging moment: ; Swing moment: ; in, The static load is the swing moment, in Nm. For swing moment dynamic load, unit: Nm; This is the static load of the swaying bending moment, in Nm; The dynamic load is the oscillation moment, unit: Nm; ω: angular velocity, unit: rad / s; t: rotation time, unit: s. Using the blade coordinate system as the reference coordinate system, the directions of centrifugal force and bending moment are defined as follows: M b The positive moment is when the paddle tip bends downwards during the swinging motion. M t : Swing moment, positive when the blade tip bends towards the leading edge (counterclockwise) / positive when it bends towards the trailing edge (clockwise); F c Centrifugal force, with blade extension being positive.

[0032] Secondly, regarding the setting of displacement boundaries, the blade root is fixed to the motor shaft with connecting bolts. Therefore, at any moment in a rotation cycle, the constraint method of the blade root is simplified to a fixed displacement boundary.

[0033] Based on the above analysis, load and displacement boundary conditions can be applied to the root of the integrated blade under both static and dynamic load conditions. (See schematic diagram below.) Figure 8 .

[0034] e. Based on the blade root force transmission path analysis and assembly relationship in the first step, establish a contact relationship between the metal cover plate and the skin, establish a constraint coupling relationship between the displacement boundary and the bolt hole, establish a constraint coupling relationship between the load application point and the skin, and establish a binding relationship between the skin and the foam and chopped fiber filler.

[0035] 4. Based on the finite element method, the strain distribution of the integrated blade root skin is calculated, and the location of strain hazard (the point of maximum strain) is identified. Based on the finite element model established in the previous step, a finite element analysis is performed on the root skin of the integrated blade. Since the skin is very thin, it can be simplified to a plane stress problem. The following mainly introduces the basic steps for strain analysis of a plane stress problem using the finite element method: a. Calculate the element stiffness matrix and the element equivalent nodal load array. The finite element method is a numerical analysis method based on the principle of minimum potential energy. In plane stress problems, the functional total potential energy π of the minimum potential energy principle is... p The matrix expression is as follows:

[0036] In the formula, ε It is a second-order strain tensor; D It is a fourth-order tensor; t It is the thickness of a two-dimensional volume; f It is a volume force acting within a two-dimensional body; T It is an area force acting on the boundary of a two-dimensional body; u It is a displacement vector; Ω is a domain (see Figure 7 ); S σ It is the boundary of the domain Ω (see Figure 9 ).

[0037] For the finite element model, the minimum potential energy principle expression for the structure should be the sum of the potential energies of each element, i.e.:

[0038] In the formula, The element stiffness matrix ; For the element equivalent nodal load array ; B This is called the strain matrix; N These are called shape functions; It is a matrix of element node displacements; Ω e It is a subdomain (see Figure 7 ); It is a subdomain Ω e The boundary (see Figure 9 ).

[0039] b. Calculate the structural stiffness matrix and the structural nodal load matrix. The element stiffness matrix and the element equivalent nodal load array obtained in step a are integrated through matrix operations to obtain the structural stiffness matrix and the structural nodal load array. The expression for calculating the structural stiffness matrix is ​​as follows: The formula for calculating the structural nodal load array is as follows: In the formula G This is the transformation matrix between the unit node degrees of freedom and the structural node degrees of freedom.

[0040] c. Calculate the structural displacement solution Through the above process, the following finite element solution equations can be obtained:

[0041] In the formula, u It is the structural displacement matrix.

[0042] d. Calculate structural strain Using the displacement solution obtained in c u The strain of each element is obtained through the formula. ε :

[0043] It should be noted that the strain near the boundary loading location may exhibit singular phenomena, thus the reliability of the boundary strain calculation results is low. However, since the loading location is not the main performance evaluation location, some elements near the loading point can be deleted to obtain a more realistic strain distribution contour map.

[0044] Using methods a through d, we obtain, as follows Figure 10 , Figure 11 The results of the radial strain calculation for the root skin of the integrated blade are shown.

[0045] 5. Analyze the lifespan of the root segment. Combining flight spectrum and load spectrum, and utilizing the strain results of the integrated blade root skin under various working conditions obtained from the above process, the life of the root section is analyzed according to the safe life analysis method.

[0046] The following is a brief description of the fatigue life analysis method for the high and low cycle zones of the full range SN curve of the root section blade skin.

[0047] The method for correcting the mean strain in high-cycle fatigue is as follows:

[0048] when R 1< R 0,

[0049] when R 1≥ R 0,

[0050] The method for correcting the mean strain of low-cycle fatigue is as follows:

[0051] in, Dynamic strain; :Static strain; : dynamic-to-static ratio 10 9 Cyclic tensile safety fatigue limit; k Fatigue reduction coefficient; k ’ Fatigue strength environmental impact coefficient; High-cycle equivalent dynamic strain; Low-cycle equivalent dynamic strain; R m-kq :Safe limit strength; K safe : Static strength safety factor.

[0052] High-frequency safety cycle count:

[0053] Low-cycle safety cycle count:

[0054] Fatigue injury:

[0055] in, High-cycle safety fatigue limit; Low-cycle safety fatigue limit; for glass cloth, α =0.1, for carbon cloth, α =0.037; N n: Fatigue failure cycles, Mc; n: Fatigue load cycles per hour, Mc; D Fatigue injury, 1 / hour; Total damage, 1 / hour; Low-cycle total damage, 1 / hour; Total high-frequency damage, 1 / hour.

[0056] Fatigue life:

[0057] in: L Safe lifespan, in hours.

Claims

1. A method for analyzing the fatigue life of an integrated blade root skin, characterized in that, This method is applied to integrated propeller blades, which include a first blade and a second blade that are centrally symmetrical and 180 degrees out of phase. The connection between the first and second blades forms a blade root, which is filled with fiber filler. The blade root has through holes that penetrate the upper and lower skins and the fiber filler. An upper metal cover plate and a lower metal cover plate are respectively installed on the upper and lower skin surfaces of the blade root. Connecting bolts are used to sequentially pass through the upper metal cover plate, the blade skin, the fiber filler, and the lower metal cover plate, directly connecting the blade root to the motor shaft. The fatigue life analysis method includes the following steps: Step 1: Analyzing the blade root... Step 1: Determine the fatigue hazard area by analyzing the load transfer path. Step 2: Extract a simplified blade root model based on the fatigue hazard area determined in Step 1. Step 3: Construct a finite element model of the blade root based on the blade root skin model in Step 2. Step 4: Calculate the strain distribution of the finite element model of the blade root using the finite element method to further identify strain hazard points. Step 5: Combine the flight spectrum and load spectrum to determine the static and dynamic load conditions. Analyze the lifespan of the root section using the static strain of the static load condition and the dynamic strain of the dynamic load condition obtained by the finite element method, following the safe life analysis method.

2. The method for analyzing the fatigue life of an integrated blade root skin according to claim 1, characterized in that: In step one, the fatigue-prone area is located in the contact area between the metal cover plate and the skin, the transition area of ​​the skin, and the skin opening area at the root of the blade; the transition area of ​​the skin refers to the area extending from the root of the blade towards the tip of the first blade and the second blade.

3. The method for analyzing the fatigue life of an integrated blade root skin according to claim 1, characterized in that: In step two, the specific process of extracting the simplified blade root model includes: Extract the model including the contact area between the metal cover plate and the skin, the skin opening area at the blade root, and the transition area of ​​the skin; The blade skin is made of multiple layers of glass cloth or carbon cloth, and the outer surface of the blade skin contour is extracted to replace the complex layered structure. The connecting bolts, upper metal cover plate, and lower metal cover plate that mate with the blade root skin are extracted directly without geometric simplification, maintaining their assembly relationship.

4. The method for analyzing the fatigue life of an integrated blade root skin according to claim 1, characterized in that: Step three specifically includes the following steps: Material parameters and constitutive definitions are defined for the blade root skin, upper metal cover plate, lower metal cover plate, and fiber filler block in the simplified blade root model. The simplified blade root model is meshed to obtain the finite element model of the blade root. Set the load and displacement boundary conditions; Configure the interaction relationships of each component in the finite element model of the blade root.

5. The method for analyzing the fatigue life of an integrated blade root skin according to claim 1, characterized in that: In step four, the strain calculation of the finite element model at the blade root is simplified to a plane stress problem, and the specific calculation is performed according to the following steps: Calculate the element stiffness matrix and the element equivalent nodal load array; Calculate the structural stiffness matrix and the structural nodal load array; Calculate the structural displacement solution; Calculate the structural strain.

6. The method for analyzing the fatigue life of an integrated blade root skin according to claim 5, characterized in that: In step four, the specific process of identifying strain hazard points based on strain distribution includes: The maximum strain point and strain red zone are obtained by statistical analysis of the structural strain field, which are used as the identified strain hazard points.

7. The method for analyzing the fatigue life of an integrated blade root skin according to claim 1, characterized in that: In step five, the safe life analysis includes high-cycle fatigue life analysis, low-cycle fatigue life analysis, and total fatigue life analysis.

8. The method for analyzing the fatigue life of an integrated blade root skin according to claim 7, characterized in that: The specific process of fatigue life analysis in the high-frequency region includes: Calculate the equivalent strain of high-cycle fatigue; Calculate the number of high-frequency safety cycles; Calculate high-cycle damage.

9. The method for analyzing the fatigue life of an integrated blade root skin according to claim 7, characterized in that: The specific process of fatigue life analysis in the low-cycle region includes: Calculate low-cycle fatigue strain; Calculate the number of high-frequency safety cycles; Calculate low-cycle damage.

10. The method for analyzing the fatigue life of an integrated blade root skin according to claim 7, characterized in that: The specific process of fatigue total life analysis includes: Calculate the total damage, including high-cycle damage and low-cycle damage; Calculate the total lifetime.

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