Method, device and medium for rapid evaluation of high-energy large-area blunt impact damage of composite aircraft structure

Abstract

The embodiment of the application provides a kind of composite aircraft structure high-energy large-area blunt impact damage rapid evaluation method, device and medium, comprising: the impact parameter and fuselage structure parameter of high-energy large-area blunt impact event are acquired;Determine the global damage mode of fuselage structure, damage location, maximum deformation and peak load;Local deformation state and local load state in the impact area of fuselage structure are calculated;The local deformation state and local load state are compared with stringer-skin interface damage criterion and gusset area damage criterion respectively, and the damage condition of the impact area is evaluated;According to the global damage mode, damage location and damage condition of the impact area, generate and output final damage assessment report.It can solve the technical problems that the evaluation speed is slow and the quantitative relationship between impact parameters and impact area damage state cannot be quickly realized in the prior art.

Description

Technical Field

[0001] This invention relates to the field of high-energy large-area blunt object impact damage assessment technology, and in particular to a rapid assessment method, device and medium for high-energy large-area blunt object impact damage of composite material aircraft structures. Background Technology

[0002] With the widespread application of carbon fiber composite materials in the main load-bearing structures of modern civil aircraft such as the Boeing 787 and Airbus A350, their impact resistance characteristics face new challenges. Accidental collisions between GSEs and aircraft structures are one of the most significant sources of damage during operation. These collisions occur at relatively low speeds, only 0.3–1.2 m / s, but due to the large mass of GSEs (2500–5500 kg) and the impact energy (100–4000 J), and the large contact area formed by the rubber buffers, this is known as High-energy wide-area blunt impact (HEWABI).

[0003] For traditional metal fuselages, HEWABI causes surface damage that is easily visually detectable. However, composite materials are characterized by being "hard and brittle," and HEWABI can cause severe, large-area damage to internal fuselage structures such as the stringer-skin interface and the corner plates connecting the bulkhead and skin, which may be almost invisible on the outside of the fuselage. This concealment of damage seriously threatens flight safety and may lead to missed detection.

[0004] Currently, the assessment of HEWABI damage mainly relies on costly experiments or computationally expensive high-precision finite element simulations, which have the following problems:

[0005] 1. It is not possible to conduct a rapid and accurate assessment at the accident site or during the design phase; the assessment process usually takes a week or even longer.

[0006] 2. It is difficult to quickly establish a quantitative relationship between impact parameters such as location, energy, and punch size and the damage status of the stringer-skin interface and corner impact areas, which restricts the development of inspection manuals and the improvement of operational efficiency. Summary of the Invention

[0007] In view of this, embodiments of the present invention provide a method, apparatus and medium for rapid assessment of high-energy large-area blunt object impact damage in composite aircraft structures, in order to solve the technical problems of slow assessment speed and inability to quickly realize the quantitative relationship between impact parameters and damage state of impact area in the prior art.

[0008] The method includes: The first aspect of this invention provides a method for rapid assessment of high-energy, large-area blunt object impact damage in composite material aircraft structures, comprising the following steps: Obtain impact parameters and fuselage structural parameters of a high-energy, large-area blunt object impact event, wherein the impact parameters include the impact location and the size of the rubber buffer; Based on the impact parameters and fuselage structure parameters, the global damage mode, damage location, maximum deformation, and peak load of the fuselage structure are determined. Based on the maximum deformation, peak load, impact location, and rubber buffer size, the local deformation state and local load state of the impact area in the fuselage structure are calculated. The local deformation state and local load state are compared with the stringer-skin interface damage criterion and the corner piece region damage criterion, respectively, to evaluate the damage status of the impact region. Based on the global damage pattern, damage location, and damage status of the impact area, a final damage assessment report is generated and output.

[0009] The second aspect of the present invention provides a rapid assessment device for high-energy large-area blunt object impact damage of composite material aircraft structures, including a memory, a processor, and a computer program stored in the memory and executable on the processor. When the processor executes the computer program, it implements the rapid assessment method for high-energy large-area blunt object impact damage of composite material aircraft structures described in the first aspect.

[0010] A third aspect of the present invention provides a computer-readable storage medium storing a computer program that executes a rapid assessment method for high-energy, large-area blunt object impact damage to composite aircraft structures as described in the first aspect.

[0011] Compared with the prior art, the beneficial effects that at least one technical solution adopted in the embodiments of this specification can achieve include at least: 1. Based on known impact scenario parameters, this solution can quickly predict the damage mode, damage location, and severity of the fuselage structure, providing efficient and reliable technical support for on-site emergency inspections, maintenance decisions, and the development of continued airworthiness documents.

[0012] 2. First, input the impact scenario parameters and use a pre-trained neural network model to quickly predict the global response of the fuselage structure; then, combine the deformation propagation and load distribution model based on the failure mechanism to analyze the local mechanical state of the impact areas such as the stringer-skin interface and corner plates; finally, apply the corresponding damage criteria to assess the local damage. Attached Figure Description

[0013] To more clearly illustrate the technical solutions of the embodiments of this application, the drawings used in the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of this application. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0014] Figure 1 This is a flowchart of a rapid assessment method for high-energy, large-area blunt object impact damage to composite material aircraft structures provided in this invention embodiment; Figure 2 This is a schematic diagram of a HEWABI impact scenario and internal damage provided in an embodiment of the present invention; Figure 3 The deformable propagation function D provided in the embodiments of the present invention prop Attenuation characteristics curve as a function of distance; Figure 4 The load allocation function F provided in this embodiment of the invention. dis A bar chart comparing the distribution results among different corner pieces; Figure 5 This is a visual schematic diagram of the four constraint terms of the physical mechanism loss function provided in the embodiment of the present invention; Figure 6 This is a schematic diagram of the structure and training objective function of the pre-trained neural network model provided in an embodiment of the present invention; Figure 7 This is a schematic diagram of a rapid assessment device for high-energy, large-area blunt object impact damage of composite material aircraft structures provided in an embodiment of the present invention. Detailed Implementation

[0015] The embodiments of this application will now be described in detail with reference to the accompanying drawings.

[0016] The following specific examples illustrate the implementation of this application. Those skilled in the art can easily understand other advantages and effects of this application from the content disclosed in this specification. Obviously, the described embodiments are only a part of the embodiments of this application, and not all of them. This application can also be implemented or applied through other different specific embodiments, and the details in this specification can also be modified or changed based on different viewpoints and applications without departing from the spirit of this application. It should be noted that, in the absence of conflict, the following embodiments and features in the embodiments can be combined with each other. Based on the embodiments in this application, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of this application.

[0017] The first aspect of this invention provides a method for rapid assessment of high-energy, large-area blunt object impact damage in composite material aircraft structures, such as... Figure 1As shown, the method includes steps S1 to S5.

[0018] Step S1: Obtain the impact parameters and fuselage structural parameters of the high-energy, large-area blunt object impact event. The impact parameters include the impact location and the size of the rubber buffer.

[0019] Specifically, the impact parameters of a high-energy, large-area blunt object impact event include the impact energy E, the characteristic length L of the rubber buffer, and the coordinate position P (x, y, z) of the impact center relative to the fuselage structure. The characteristic length L of the rubber buffer is used to simulate the contact length of the GSE buffer, and the coordinate position P of the impact center relative to the fuselage structure is used to determine the relative position with respect to the bulkhead, stringer, and fuselage bottom. A schematic diagram of the HEWABI impact scenario is shown below. Figure 2 As shown.

[0020] The fuselage structural parameters specifically refer to the basic configuration parameters of the target fuselage frame structure, such as the frame spacing d. f , stringer spacing d s And skin thickness t, etc.

[0021] Step S2: Determine the global damage mode, damage location, maximum deformation, and peak load of the fuselage structure based on the impact parameters and fuselage structure parameters.

[0022] This step is completed based on a pre-trained neural network model. The structure of the neural network model is as follows: Figure 6 As shown, it includes an 8-dimensional input layer, three fully connected hidden layers, and a 4-dimensional output layer. The input variables of the input layer are the impact energy E, the impact position coordinates x, y, z, the buffer length L, and the frame spacing d. f , stringer spacing d s The skin thickness is t. The number of neurons in the fully connected hidden layers are 128, 256, and 128, respectively, with ReLU activation function. Dropout regularization is applied after the second hidden layer, with a dropout rate of 0.2. The output layer outputs global damage pattern classification, damage location coordinates (x, y, z), and maximum deformation D. max and peak load F max .

[0023] During neural network model training, high-energy, large-area blunt object impact simulation data are used to train the neural network model. The optimization objective function for training is a weighted sum of data prediction loss and physical mechanism loss, i.e.: Min Loss = Loss data + λ·Loss mech In the formula, λ is the weighting coefficient, ranging from 0.1 to 0.5, and the optimal value is determined through cross-validation; Loss dataTo predict the loss for the data, a weighted sum of the mean squared error of the output variables and the classification cross-entropy is used; Loss mech To address the physical mechanism loss, a regularization constraint is applied to the physical consistency of the prediction results based on four physical constraints: energy conservation, stiffness consistency, geometric feasibility, and strength limit. During neural network training, the network's weight parameters are continuously adjusted to minimize this Min Loss value. The smaller the Min Loss value, the closer the prediction results are to reality and the more they conform to physical laws.

[0024] The training termination condition is as follows: when the relative decrease in the loss function value over K consecutive training epochs is lower than a preset threshold ε, the training is considered to have converged and training is stopped; simultaneously, a maximum training epoch N is set. max An upper bound is set to prevent infinite loops when training fails to converge. That is: Convergence criterion: |Loss(t)-Loss(tK)| / Loss(tK)<ε In the formula: Loss(t) is the loss function value in the t-th training epoch; K is the comparison window for the rate of decrease, with a recommended value of 15 to 20 epochs; ε is the convergence threshold, with a recommended value of 8.10. -4 ~10 - ³;N max For the maximum number of training rounds, a value of 800 to 1000 is recommended.

[0025] When the convergence criterion is met or the number of training rounds reaches N max When the training stops, the model parameters with the minimum loss on the validation set are taken as the final model.

[0026] The data prediction loss Loss data The specific expression is: Loss data =(D max -D max_true )²+(F max -F max_true )²+[(xx true )²+(yy true )²+(zz true )²]-Loss class , D max F max x, y, z, and mode(k) are the prediction results of the neural network; D max_true F max_true x true y true z true The actual values ​​provided for the simulation data.

[0027] Lossclass The loss value is used to classify damage patterns. Based on the training results of the input sample data, the neural network automatically outputs a probability value for each possible damage pattern under the predicted working condition. This value measures the difference between the predicted probability distribution and the actual damage pattern (sample data). The more accurate the prediction, the smaller the value.

[0028] The physical mechanism loss Loss mech The specific expression is: Loss mech =w1·[max(0,k eff ·D max ² / (2·E impact )-1)]²+w2·[max(0,|F max / (k eff ·D max )-1|-0.3)]²+w3·[max(0,D max / (0.3·d frame )-1)]²+w4·[max(0,F max / (L·100·σ_ crush )-0.8)]², k eff =E skin ·t·L / d frame , In the formula: k eff E represents the effective stiffness of the fuselage structure, expressed in N / mm. skin t is the elastic modulus of the skin, in MPa; t is the skin thickness, in mm; L is the characteristic length of the rubber buffer, in mm; d frame This refers to the spacing between the partitions, in mm; D max E represents the maximum deformation predicted by the neural network, in mm. impact The input impact energy is expressed in J or F. max The peak load predicted by the neural network is in N; σ crush The external crush strength of the skin is measured in MPa and is determined by a quasi-static indentation test. w1, w2, w3, and w4 are the weighting coefficients of each constraint item, with recommended values ​​of 1.0, 2.0, 2.0, and 1.0, respectively.

[0029] The physical meaning and feasible region of the four constraints are as follows: Figure 5 As shown. Figure 5 (a) Given the energy conservation constraint, the predicted strain energy U strain No more than the input impact energy E impact U strain =k eff ·D max ² / 2; Figure 5(b) For stiffness consistency constraints, the predicted peak load F max With maximum deformation D max It satisfies an approximately linear relationship, allowing a deviation range of ±30%; Figure 5 (c) represents the geometric feasibility constraint, with the maximum deformation D. max No more than the spacing d between the frames frame 30%; Figure 5 (d) represents the ultimate strength constraint, where the impact contact stress does not exceed the skin surface crush strength σ. crush 80%.

[0030] After training, the impact parameters and fuselage structure parameters can be directly input into the pre-trained neural network model to quickly obtain the global damage pattern, damage location, maximum deformation, and peak load of the fuselage structure.

[0031] The maximum deformation is the maximum deformation D of the fuselage impact area. max。

[0032] Step S3: Based on the maximum deformation, peak load, impact location, and rubber buffer size, calculate the local deformation state and local load state of the impact area in the fuselage structure.

[0033] This step is based on D obtained in step S2. max and F max Combining the impact position P and the punch length L from step S1, the local state is analyzed using the established mechanism model.

[0034] Specifically, this step can be based on deformation propagation analysis using a deformation propagation function to obtain the local deformation state of the impact region in the fuselage structure.

[0035] Load distribution analysis is performed based on the load distribution function to calculate the local load state of the impact region in the fuselage structure.

[0036] The deformation propagation function is: D local =D max ·D prop , Among them, D prop The deformation propagation coefficient is calculated using the following formula based on the normalized distance between the impact point and the frame / girder, the dimensions of the rubber buffer, and the thickness of the skin: D prop =exp(-β1·d frame )·exp(-β2·d stringer )·(L / L ref )^(-α)·(t ref / t)^γ, In the formula, Dlocal D is in a state of local deformation. max For maximum deformation, d frame d is the axial distance from the impact center to the nearest partition, in mm. stringer β1 is the circumferential distance from the impact center to the nearest stringer, in mm; β1 is the axial attenuation coefficient, pre-calibrated through material mechanics tests, with a value ranging from 0.002 to 0.008 mm. - ¹; β2 is the circumferential attenuation coefficient, pre-calibrated through material mechanics tests, with a value ranging from 0.001 to 0.005 mm. - ¹; L is the characteristic contact length of the rubber buffer, in mm; L ref The reference length is 200 mm; α is the size effect index, determined through parameter fitting, with a value range of 0.3 to 0.4; t is the skin thickness in mm; t ref The reference thickness is 2.5 mm; γ is the thickness correction index, which is determined by parameter fitting and has a value range of 1.2 to 1.8.

[0037] The load distribution function is: F i =F max ·F dis i , In the formula, F i For the load borne by the i-th corner piece, F max For peak load; F dis i This is the load distribution coefficient for the i-th corner piece, with a value ranging from 0 to 1. Specifically, F dis i =(K i ·e^(-d i / d char )) / Σ j (K j ·e^(-d j / d char )), K i =(E clip ·A clip ) / h clip , d i =√[(x i -x impact )²+(y i -y impact )²+(2·|z i -z impact |)²], d char =L+2t·tanθ, In the formula, K i E represents the equivalent compressive stiffness of the i-th corner piece, expressed in N / mm, where E clip • The elastic modulus of the corner piece material, in MPa; A clip The cross-sectional area of ​​the corner piece is measured in mm², where A clip =b×t clip b is the width of the corner piece measured along the circumferential direction of the skin, in mm, t clip h represents the thickness of the corner piece in mm. clip d represents the height of the corner piece, in mm. i x represents the equivalent spatial distance from the i-th corner piece to the impact center, in mm. i y i z i Let x be the position coordinate of the i-th corner piece. impact y imapct z impact The coordinates of the impact point's center are: x is along the fuselage axis (longitudinal direction, i.e., along the bulkhead distribution direction); y is along the fuselage circumferential direction (circumferential direction, i.e., along the skin circumference); z is along the fuselage radial direction (normal direction, i.e., along the skin thickness direction); d char t is the stress diffusion characteristic length in mm, L is the characteristic contact length of the rubber buffer in mm, t is the skin thickness in mm, and θ is the stress diffusion half-angle of the composite material, ranging from 30° to 45°.

[0038] Example of load distribution function distribution results between different corner pieces: Figure 4 As shown. Figure 4 The distance d of the five corner pieces, namely corner pieces 1 to 5, in a given impact scenario is shown. i With load distribution factor F dis i In the columnar comparison, corner piece 2 is closest to the impact center, specifically, d2=60mm, and its load distribution factor is the largest, specifically, F dis 2 =0.243; Corner piece 5 is the furthest away, specifically, d5=200mm, and its load distribution factor is the smallest, specifically, F dis 5 =0.142, which conforms to the physical law of stress attenuation with distance in Saint-Venant's principle.

[0039] Through F dis i After the function determines the load distribution coefficients for each corner piece, the compressive load Fc, bending moment Mb, and torque Tt borne by the i-th corner piece are further decomposed using the following formula: F c =F i ×cosθz M b =0.15×F i ×h clip T t =0.10×F i ×t clip Where, θ z θ is the angle between the direction of the impact force and the axial direction of the corner piece, i.e., the normal direction of the skin. When the distance between the impact point and the corner piece is relatively close and the influence of the fuselage curvature is negligible, θ z ≈0°, cosθ z ≈1, at this time F c ≈F i The proportionality coefficients 0.15 and 0.10 are empirical values ​​obtained through finite element parametric analysis and can be calibrated according to the specific machine structure.

[0040] Step S4: Compare the local deformation state and local load state with the stringer-skin interface damage criterion and the corner piece region damage criterion, respectively, to evaluate the damage status of the impact region.

[0041] The stringer-skin interface damage criterion and the corner plate area damage criterion are pre-established damage criterion models. The damage status of the impact area includes whether debonding occurs at the impact area interface, the severity of debonding, whether the corner plate fails, the specific failure mode, and the failure sequence.

[0042] Torch-skin interface damage assessment: The local deformation state at the interface is determined by D. local Substitute the following damage criterion based on the BK (Benzeggagh-Kenane) mixed-mode fracture criterion to determine whether debonding occurs at the interface and the severity of the debonding.

[0043] The damage criterion for the stringer-skin interface is as follows: DI interface =[E skin ·D² local / (2t)+G skin ·t·(0.3·D local / t)² / 2] / [GI c +(GII c -GI c )·β^η] in: β=G skin ·(0.3·D local / t)² / [E skin ·(D local / t)²+G skin ·(0.3·Dlocal / t)²] Where: DI interface E represents the interface damage index. skin G is the equivalent elastic modulus of the skin, in MPa; skin D is the skin shear modulus, in MPa. local The local deformation at the interface is represented by t, in mm; t is the skin thickness, in mm; GI c Type I interlaminar fracture toughness, measured in N / mm, determined by a Type I interlaminar fracture toughness test; GII c Type II interlaminar fracture toughness, in N / mm, is determined by a Type II interlaminar fracture toughness test; η is the BK mixed mode parameter, calibrated by a Type I / II mixed mode test, with a value range of 1.5 to 3.0; β is the proportion of Type II energy release rate in the total energy release rate.

[0044] The grading criteria for the severity of debonding are shown in Table 1.

[0045] Table 1. Criteria for Damage at the Truss-Skin Interface

[0046] The local load state at the corner piece is compared with the damage criteria of the corner piece area to determine whether the corner piece has failed, the specific failure mode, and the failure sequence.

[0047] Corner plate area damage assessment: Apply the compressive load F at the corner plate. c Bending moment M b Torque T t Substitute the three local load states into the following combined load interaction failure criterion to determine whether the corner piece has failed, the specific failure mode, and the failure sequence.

[0048] The criterion for damage to the corner patch area is: FI combined =[(|F c / A clip +M b ·y max / I| / σ 11_C )²+(|T t / J| / τ 13_S ) α ]^(1 / α) Where: FI combined F represents the damage index of the corner piece combination. c The compressive load borne by the corner piece, in NF. c =F i ×cosθ z , where θ zIt is the angle between the direction of the impact force and the axial direction of the corner piece, i.e., the normal direction of the skin. clip M represents the cross-sectional area of ​​the corner piece, in mm². b The bending moment borne by the corner piece, in N·mm, M b =0.15·F i ·h clip ;y max I is the distance from the neutral axis of the corner section to the outermost edge, in mm; I is the moment of inertia of the corner section, in mm. 4 ;σ 11_C T represents the fiber-direction compressive strength of the composite material, expressed in MPa. t The torque borne by the corner piece, expressed in N·mm, T t =0.10·F i ·t clip J is the torsional constant of the corner plate section, in mm. 4 ;τ 13_S α represents the interlaminar shear strength of the composite material, in MPa; α is the interaction exponent, with a value of 2.0.

[0049] The classification criteria for the degree of damage to the corner piece area are based on the corner piece combination damage index, as shown in Table 2.

[0050] Table 2 Criteria for Corner Plate Area Damage Judgment

[0051] The failure mode determination method is as follows: First, calculate the fiber compression failure index, matrix failure index, and interlaminar shear failure index separately. The fiber compression failure index is: FI fc =|F c / A clip +M b ·y max / I| / σ 11_C , The matrix failure index is: FI mt =(T t / J / σ 22_T )², The interlaminar shear failure index: FI is =|0.3·F c / A+0.7·T t / J| / τ 13_S， In the formula, σ 22_T This represents the transverse tensile strength of the composite material, expressed in MPa.

[0052] The fiber compression failure index, matrix failure index, and interlaminar shear failure index are then sorted from largest to smallest to obtain the predicted failure sequence and the dominant failure mode under the current load condition. The dominant failure mode under the current load condition is the index corresponding to the maximum value.

[0053] Local load states include compression, bending, and torsion; specific failure modes include chamfering delamination and connection zone failure.

[0054] Step S5: Generate and output the final damage assessment report based on the global damage pattern, damage location, and damage status of the impact area.

[0055] Integrating the global prediction from step S2 with the local assessment results from step S4, a comprehensive damage assessment report is generated. The report is output in a visual format, marking the predicted damage areas, damage patterns, and severity levels on a schematic diagram of the fuselage structure, and providing recommended key inspection areas and suggested inspection methods.

[0056] This invention combines the efficiency of data-driven approaches with the accuracy of physical mechanisms, enabling rapid and reliable assessment of HEWABI, a type of damage that is concealed and severe. It provides an efficient technical means for post-accident emergency inspections, maintenance decisions, and continued airworthiness assurance of aircraft.

[0057] This invention achieves the following technical effects through the aforementioned mathematical functions: The deformation propagation function integrates four physical mechanisms—axial attenuation, circumferential attenuation, size effect, and thickness correction—and uses different attenuation coefficients in the axial and circumferential directions to reflect the anisotropy of composite material stiffness, providing better extrapolation and prediction capabilities for new impact locations not included in the training; The load distribution function simultaneously incorporates corner plate stiffness and distance weighting, and calculates the three load components of compression, bending, and torsion respectively, achieving a complete description of complex stress states; The stringer-skin interface damage criterion adopts the BK hybrid mode fracture criterion, considering the mixed effect of Type I and Type II fractures; The corner plate region damage criterion integrates the in-plane criterion and the interlaminar shear criterion, and through interactive exponential coupling of the two failure mechanisms, it can determine four failure modes and their order of occurrence at once; The physical mechanism loss function introduces four constraints—energy conservation, stiffness consistency, geometric feasibility, and strength limit—as regularization terms into the training objective, making the prediction results automatically satisfy physical laws, significantly reducing unreliable predictions in sparse data regions, and improving the model's robustness and engineering applicability.

[0058] The second aspect of this invention provides a rapid assessment device for high-energy, large-area blunt object impact damage in composite material aircraft structures, such as... Figure 7As shown, it includes a memory 201, a processor 202, and a computer program stored in the memory and executable on the processor. When the processor executes the computer program, it implements a rapid assessment method for high-energy, large-area blunt object impact damage of any of the composite material aircraft structures described in the first aspect above.

[0059] A third aspect of the present invention provides a computer-readable storage medium storing a computer program that executes any of the rapid assessment methods for high-energy, large-area blunt object impact damage of composite aircraft structures according to the first aspect.

[0060] Specifically, computer-readable storage media, including both permanent and non-permanent, removable and non-removable media, can store information using any method or technology. Information can be computer-readable instructions, data structures, program modules, or other data. Examples of computer-readable storage media include, but are not limited to, phase-change memory (PRAM), static random access memory (SRAM), dynamic random access memory (DRAM), other types of random access memory (RAM), read-only memory (ROM), electrically erasable programmable read-only memory (EEPROM), flash memory or other memory technologies, CD-ROM, digital versatile optical disc (DVD) or other optical storage, magnetic tape, magnetic disk storage or other magnetic storage devices, or any other non-transferable medium that can be used to store information accessible by a computing device. As defined herein, computer-readable storage media does not include transient media, such as modulated data signals and carrier waves.

[0061] The above description is merely a preferred embodiment of the present invention and is not intended to limit the present invention. For those skilled in the art, various modifications and variations can be made to the embodiments of the present invention. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the protection scope of the present invention.

Claims

1. A rapid assessment method for high-energy, large-area blunt object impact damage in composite material aircraft structures, characterized in that, Includes the following steps: Obtain impact parameters and fuselage structural parameters of a high-energy, large-area blunt object impact event, wherein the impact parameters include the impact location and the size of the rubber buffer; Based on the impact parameters and the fuselage structure parameters, the global damage mode, damage location, maximum deformation, and peak load of the fuselage structure are determined. Based on the maximum deformation, the peak load, the impact location, and the size of the rubber buffer, the local deformation state and local load state of the impact area in the fuselage structure are calculated. The local deformation state and local load state are compared with the stringer-skin interface damage criterion and the corner piece region damage criterion, respectively, to evaluate the damage status of the impact region. Based on the global damage pattern, damage location, and damage status of the impact area, a final damage assessment report is generated and output.

2. The rapid assessment method for high-energy, large-area blunt object impact damage to composite aircraft structures as described in claim 1, characterized in that, The process of determining the global damage mode, damage location, maximum deformation, and peak load of the fuselage structure based on the impact parameters and fuselage structural parameters includes: The impact parameters and fuselage structure parameters are input into a pre-trained neural network model to obtain the global damage pattern, damage location, maximum deformation, and peak load of the fuselage structure.

3. The rapid assessment method for high-energy, large-area blunt object impact damage to composite aircraft structures as described in claim 2, characterized in that, The step of inputting the impact parameters and fuselage structure parameters into a pre-trained neural network model also includes: The neural network model was trained using high-energy, large-area blunt object impact simulation data. The optimization objective function of the training was a weighted sum of data prediction loss and physical mechanism loss.

4. The rapid assessment method for high-energy, large-area blunt object impact damage to composite aircraft structures as described in claim 1, characterized in that, Based on the maximum deformation, peak load, impact location, and rubber buffer dimensions, the local deformation and local load states of the impact region in the fuselage structure are calculated, including: Based on the maximum deformation, peak load, impact location, and rubber buffer dimensions, the local deformation state of the impact region in the fuselage structure is calculated using the deformation propagation function, and the local load state of the impact region in the fuselage structure is calculated using the load distribution function.

5. The rapid assessment method for high-energy, large-area blunt object impact damage to composite aircraft structures as described in claim 4, characterized in that, The deformation propagation function is: D local =D max ·D prop , in, D prop =exp(-β1·d frame )·exp(-β2·d stringer )·(L / L ref )^(-α)·(t ref / t)^γ, In the formula, D local D is in a state of local deformation. max For maximum deformation, D prop d is the deformation propagation coefficient, exp is the exponential function, β1 is the axial attenuation coefficient, and d frame d is the axial distance from the impact center to the nearest partition, β2 is the circumferential attenuation coefficient, and d is the axial distance from the impact center to the nearest partition. stringer L is the circumferential distance from the impact center to the nearest stringer, and L is the characteristic contact length of the rubber buffer. ref The reference length is α, the size effect exponent is t. ref The reference thickness is t, where t is the skin thickness and γ is the thickness correction index.

6. The rapid assessment method for high-energy, large-area blunt object impact damage to composite material aircraft structures as described in claim 4, characterized in that... The load allocation function is: F i =F max ·F dis i , in, F dis i =(K i ·e^(-d i / d char )) / Σ j (K j ·e^(-d j / d char )), K i =(E clip ·A clip ) / h clip , d i =√[(x i -x impact )²+(y i -y impact )²+(2·|z i -z impact |)²], d char =L+2t·tanθ, In the formula, F i For the load borne by the i-th corner piece, F max For peak load, F dis i K is the load distribution factor for the i-th corner piece. i Let be the equivalent compressive stiffness of the i-th corner piece, e be the base of the natural logarithm, and d be the equivalent compressive stiffness of the i-th corner piece. i Let d be the equivalent spatial distance from the i-th corner piece to the impact center. char E is the characteristic length of stress diffusion. clip Let A be the elastic modulus of the corner piece material. clip h is the cross-sectional area of ​​the corner piece. clip x is the height of the corner piece. i Let y be the x-coordinate of the position of the i-th corner piece. i Let z be the ordinate of the position of the i-th corner piece. i Let x be the vertical coordinate of the position of the i-th corner piece. impact Let y be the x-coordinate of the center point of the impact location. imapct Let z be the ordinate of the center point of the impact location. impact θ is the vertical coordinate of the center point of the impact position, L is the characteristic contact length of the rubber buffer, t is the skin thickness, and θ is the stress diffusion half angle of the composite material.

7. The rapid assessment method for high-energy, large-area blunt object impact damage to composite aircraft structures as described in claim 1, characterized in that, The step of comparing the local deformation state and local load state with the stringer-skin interface damage criterion and the corner piece region damage criterion, respectively, to evaluate the damage status of the impact region includes: The local deformation state at the interface is compared with the stringer-skin interface damage criterion to determine whether debonding has occurred at the interface and the severity of the debonding. The damage criterion for the stringer-skin interface is as follows: This interface =[E skin ·D² local / (2t)+G skin ·t·(0.3·D local / t)² / 2] / [GI c +(GII c -GI c )·β^η] in: β=G skin ·(0.3·D local / t)² / [E skin ·(D local / t)²+G skin ·(0.3·D local / t)²], Where: DI interface E represents the interface damage index. skin D is the equivalent elastic modulus of the skin. local The local deformation state at the interface is represented by t, where t is the skin thickness and G is the local deformation state at the interface. skin D is the skin shear modulus. local GI represents the local deformation state at the interface. c For type I interlaminar fracture toughness, GII c η represents the type II interlaminar fracture toughness, η is the BK mixed mode parameter, and β is the proportion of type II energy release rate in the total energy release rate. The local load state at the corner piece is compared with the damage criteria of the corner piece area to determine whether the corner piece has failed, the specific failure mode, and the failure sequence. The criterion for damage to the corner patch area is: FI combined =[(|F c / HAS clip +M b ·y max / I| / σ 11_C )²+(|T t / J| / τ 13_S ) α ]^(1 / α), Where: FI combined F represents the damage index of the corner piece combination. c For the compressive load borne by the corner piece, A clip M is the cross-sectional area of ​​the corner piece. b The bending moment y is borne by the corner piece. max Let σ be the distance from the neutral axis of the corner section to the outermost edge, I be the moment of inertia of the corner section, and σ be the distance from the neutral axis to the outermost edge. 11_C T represents the fiber-direction compressive strength of the composite material. t Let J be the torque borne by the corner piece, J be the torsional constant of the corner piece cross section, and τ be the torque borne by the corner piece. 13_S α represents the interlaminar shear strength of the composite material, and α is the interaction exponent.

8. The rapid assessment method for high-energy, large-area blunt object impact damage to composite material aircraft structures as described in claim 7, characterized in that, The failure mode determination method is as follows: Calculate the fiber compression failure index, matrix failure index, and interlaminar shear failure index separately. The fiber compression failure index FI fc for: FI fc =|F c / HAS clip +M b ·y max / I| / σ 11_C , The matrix failure index FI mt for: FI mt =(T t / J / σ 22_T )², The interlaminar shear failure index FI is for: FI is =|0.3·F c / A+0.7·T t / J| / τ 13_S , In the formula, F c For compressive load, M b For bending moment, T t For torque, σ 22_T The transverse tensile strength of the composite material; The fiber compression failure index, matrix failure index, and interlaminar shear failure index are sorted from largest to smallest to obtain the predicted failure sequence and the dominant failure mode under the current load state. The dominant failure mode under the current load state is the index corresponding to the maximum value.

9. A rapid assessment device for high-energy, large-area blunt object impact damage in composite material aircraft structures, comprising a memory, a processor, and a computer program stored in the memory and executable on the processor, characterized in that, When the processor executes the computer program, it implements a rapid assessment method for high-energy, large-area blunt object impact damage of composite material aircraft structures according to any one of claims 1 to 7.

10. A computer-readable storage medium, characterized in that, The computer-readable storage medium stores a computer program that executes a rapid assessment method for high-energy, large-area blunt object impact damage of a composite material aircraft structure according to any one of claims 1 to 7.

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