Method, device and equipment for acquiring shape tolerance of aircraft component assembly, and storage medium

Abstract

The application discloses an aircraft component assembly shape tolerance acquisition method, device, equipment and storage medium, based on the real assembly relationship of the aircraft, considering the part manufacturing error and assembly error, adopting a three-dimensional tolerance modeling method for prediction, avoiding the condition of relying on experience for judgment, improving the accuracy of adjustment; and, adopting a simulation calculation method to predict the aircraft component shape, getting rid of the need for measuring and judging the real aircraft shape, low cost and high efficiency; further, the existing manufacturing tolerance design error model can be used to predict the aircraft shape error, and the manufacturing process tolerance can also be designed based on the target aircraft shape error requirement, avoiding the process of repeated trial and error and adjustment of the traditional method, and quality and efficiency are considered.

Description

Technical Field

[0001] This application relates to the field of aircraft assembly, and in particular to a method, apparatus, equipment and storage medium for obtaining external appearance differences in aircraft component assembly. Background Technology

[0002] The accuracy of an aircraft's shape is crucial to its aerodynamic characteristics, significantly impacting its flight performance and quality. Both aircraft designers and manufacturers strive for an actual shape that closely resembles the theoretical shape. However, unavoidable factors such as manufacturing errors, assembly errors, and deformation due to aircraft weight inevitably result in some deviation between the actual produced aircraft's shape and the theoretical shape.

[0003] In aircraft manufacturing, process engineers pre-design and allocate manufacturing tolerances for parts and assembly tolerances for components. They then inspect the aircraft's external shape using inspection fixtures or digital measurement methods (such as laser trackers), adjust tolerances based on the inspection results, and use these tolerances to adjust the aircraft's external shape to ensure the accuracy of the aircraft components' shapes. However, currently, the adjustment of tolerances lacks theoretical basis and mainly relies on the work experience of process engineers, resulting in relatively poor accuracy. Summary of the Invention

[0004] The main objective of this application is to provide a method, apparatus, equipment, and storage medium for obtaining the external tolerance of aircraft component assembly, aiming to solve the technical problem of poor accuracy in existing tolerance adjustment methods.

[0005] To achieve the above objectives, this application provides a method for obtaining the external appearance difference of aircraft component assembly, comprising:

[0006] Obtain process assembly information for the analysis area of ​​the shape error of the target aircraft component; wherein, the process assembly information includes assembly tolerances;

[0007] Based on the process assembly information, the assembly relationship chain of the analysis area is obtained; wherein, the assembly relationship chain is the aircraft component assembly relationship from the analysis datum to the analysis area;

[0008] Based on the assembly relationship chain, an error calculation model is constructed;

[0009] Based on the error calculation model, the predicted index of the shape error of the target aircraft component is obtained;

[0010] Based on the predicted indicators, the assembly appearance difference of the target aircraft component is obtained.

[0011] Optionally, the step of constructing an error calculation model based on the assembly relationship chain includes:

[0012] Based on the assembly relationship chain, obtain the first point vector and the first normal vector of any point in the analysis region in the reference coordinate system, as well as the second point vector and the second normal vector of that point in the analysis region coordinate system; wherein, the Z-axis of the analysis region coordinate system is in the same direction as the first normal vector;

[0013] Based on the first point vector, the first normal vector, the second point vector, and the second normal vector, obtain the coordinate system transformation relationship under the theoretical numerical model state;

[0014] Based on the coordinate system transformation relationship, standard calculation models for aircraft component shapes and error calculation models for aircraft component shapes are constructed respectively.

[0015] The error calculation model is constructed based on the standard calculation model of the aircraft component shape and the error calculation model of the aircraft component shape.

[0016] Optionally, the step of obtaining the coordinate system transformation relationship under the theoretical numerical model state based on the first point vector, the first normal vector, the second point vector, and the second normal vector includes:

[0017] The coordinate system transformation relationship is obtained through the following formula:

[0018]

[0019]

[0020] in, Let be the vector of the first point, and its coordinates are... Let the second point vector have coordinates of... Let be the first normal vector, and its coordinates are... Let be the first normal vector, and its coordinates are... The homogeneous matrix represents the transformation relationship from coordinate system C1 to coordinate system C0. The homogeneous matrix represents the transformation relationship from coordinate system C2 to coordinate system C1. The homogeneous matrix represents the transformation relationship from coordinate system C3 to coordinate system C2. C represents n coordinate system to C n-1 The homogeneous matrix representing the transformation relationship between coordinate systems.

[0021] Optionally, the step of constructing the standard calculation model of the aircraft component shape and the calculation model of the aircraft component shape error based on the coordinate system transformation relationship includes:

[0022] The standard calculation model for the shape of the aircraft component is constructed using the following relationship:

[0023]

[0024] Where, ε i The theoretical thickness of the outer skin representing the frame of an aircraft component. The homogeneous matrix represents the transformation relationship from coordinate system C1 to coordinate system C0. The homogeneous matrix represents the transformation relationship from coordinate system C2 to coordinate system C1. The homogeneous matrix represents the transformation relationship from coordinate system C3 to coordinate system C2. C represents n coordinate system to C n-1 Homogeneous matrix relating coordinate system transformations;

[0025] The following relationship is used to construct a calculation model for the shape error of the aircraft component:

[0026]

[0027] Where, ε′ i The actual thickness (m layers) of the skin (representing the outer frame of an aircraft component), Δδ n represent Error matrix The homogeneous matrix represents the transformation relationship from coordinate system C1 to coordinate system C0. The homogeneous matrix represents the transformation relationship from coordinate system C2 to coordinate system C1. The homogeneous matrix represents the transformation relationship from coordinate system C3 to coordinate system C2. C represents n coordinate system to C n-1 Homogeneous matrix relating coordinate system transformations;

[0028] The step of constructing the error calculation model based on the aircraft component shape standard calculation model and the aircraft component shape error calculation model includes:

[0029] The error calculation model is constructed using the following relationship:

[0030]

[0031] Among them, M i For the calculation model of aircraft component shape error, N i A standard calculation model for the shape of aircraft components.

[0032] Optionally, the step of obtaining the predicted index of the shape error of the target aircraft component based on the error calculation model includes:

[0033] Based on the assembly tolerances, obtain the error matrix;

[0034] Obtain the actual thickness of the outer skin of the aircraft component frame;

[0035] Based on the error matrix and the actual thickness of the skin, the aircraft component shape error calculation model is simulated to obtain the aircraft component shape error simulation dataset.

[0036] Based on the aircraft component shape error simulation dataset and the error calculation model, an error calculation dataset is obtained;

[0037] The prediction index is obtained based on the error calculation dataset; wherein the prediction index is the mean and standard deviation of the data in the error calculation dataset.

[0038] Optionally, the step of obtaining the assembly appearance difference of the target aircraft component based on the predicted index includes:

[0039] Determine whether the manufacturing tolerance pass rate of the assembly tolerance is greater than or equal to meeting the preset process requirements; wherein, the manufacturing tolerance pass rate is obtained based on the prediction index;

[0040] If so, the assembly tolerance is the difference in the external appearance of the assembly;

[0041] If not, the assembly tolerance corresponding to the error parameter with the largest contribution in the error matrix is ​​selected for adjustment to obtain the assembly external tolerance value.

[0042] Optionally, the manufacturing tolerance pass rate can be obtained through the following relationship:

[0043]

[0044] Among them, T max For the upper limit of the shape error of aircraft components, T min Here, μ and σ represent the lower limit of the shape error of aircraft components, and μ is the prediction index. σ is the standard deviation.

[0045] Optionally, the contribution can be obtained through the following relationship:

[0046]

[0047] Among them, t i For assembly tolerances, σ is the standard deviation.

[0048] If not, the step of selecting the assembly tolerance corresponding to the error parameter with the largest contribution in the error matrix for adjustment to obtain the assembly external tolerance value includes:

[0049] Modify the tolerance zone position and tolerance zone width of the assembly tolerance corresponding to the error parameter with the largest contribution to obtain the assembly external tolerance value.

[0050] Furthermore, to achieve the above objectives, this application also provides an apparatus for obtaining external shape differences in aircraft component assembly, comprising:

[0051] The process assembly information acquisition module is used to acquire process assembly information of the analysis area of ​​the shape error of the target aircraft component; wherein, the process assembly information includes assembly tolerances.

[0052] The assembly relationship chain acquisition module is used to acquire the assembly relationship chain of the analysis area based on the process assembly information; wherein, the assembly relationship chain is the aircraft component assembly relationship from the analysis datum to the analysis area;

[0053] An error calculation model construction module is used to construct an error calculation model based on the assembly relationship chain;

[0054] The prediction index acquisition module is used to acquire the prediction index of the shape error of the target aircraft component according to the error calculation model.

[0055] The tolerance value acquisition module is used to acquire the assembly external tolerance value of the target aircraft component based on the prediction index.

[0056] In addition, to achieve the above objectives, this application also provides a computer device including a memory and a processor, wherein the memory stores a computer program and the processor executes the computer program to implement the above-described method.

[0057] In addition, to achieve the above objectives, this application also provides a computer-readable storage medium storing a computer program, on which a processor executes the computer program to implement the above-described method.

[0058] The beneficial effects that this application can achieve.

[0059] This application proposes a method, apparatus, device, and storage medium for obtaining the external shape error of an aircraft component assembly. The method involves acquiring process assembly information of the analysis area of ​​the external shape error of a target aircraft component; wherein the process assembly information includes assembly tolerances; obtaining an assembly relationship chain of the analysis area based on the process assembly information; wherein the assembly relationship chain is the aircraft component assembly relationship from the analysis datum to the analysis area; constructing an error calculation model based on the assembly relationship chain; obtaining a prediction index of the external shape error of the target aircraft component based on the error calculation model; and obtaining the assembly external shape error value of the target aircraft component based on the prediction index. Based on the actual assembly relationships of the aircraft, and considering the manufacturing and assembly errors of parts, a three-dimensional tolerance modeling method is used for prediction, avoiding reliance on experience for judgment and improving the accuracy of adjustments. Furthermore, simulation calculation methods are used to predict the shape of aircraft components, eliminating the need for tooling that requires measuring and judging the actual shape of the aircraft, resulting in low cost and high efficiency. Moreover, it can predict aircraft shape errors based on existing manufacturing tolerance design error models, or it can design process tolerances for manufacturing process tolerances based on the target aircraft shape error requirements, avoiding the repeated trial and error and adjustment process of traditional methods, thus balancing quality and efficiency. Attached Figure Description

[0060] Figure 1 This is a schematic diagram of the computer device structure for the hardware operating environment involved in the embodiments of this application;

[0061] Figure 2 A flowchart illustrating a method for obtaining external appearance differences in aircraft component assembly, provided as an embodiment of this application;

[0062] Figure 3 A schematic diagram of the functional modules of an aircraft component assembly external appearance difference acquisition device provided in an embodiment of this application;

[0063] Figure 4 A schematic diagram illustrating the principle of a method for obtaining external appearance differences in aircraft component assembly, provided in an embodiment of this application;

[0064] Figure 5 A schematic diagram illustrating the principle of a method for obtaining external appearance differences in aircraft component assembly, provided in an embodiment of this application;

[0065] Figure 6 An assembly relationship transfer chain diagram for an aircraft component assembly external appearance difference acquisition method provided in this application embodiment;

[0066] Figure 7 This application provides a schematic diagram illustrating the geometric feature position coordinate transformation in different coordinate systems for a method of obtaining external shape differences in aircraft component assembly, as provided in an embodiment of the present application.

[0067] Figure 8 This is a schematic diagram of planar shape and position features for a method of obtaining external shape differences in aircraft component assembly, provided in an embodiment of this application.

[0068] Figure 9 A schematic diagram of cylindrical shape and position features for a method of obtaining external shape difference in aircraft component assembly provided in an embodiment of this application;

[0069] Figure 10 This application provides an embodiment of a method for obtaining external appearance differences in aircraft component assembly, illustrated by an exploded view of a typical aircraft component.

[0070] Figure 11 This is a schematic diagram of the coordinate system of a typical aircraft component, which is provided in an embodiment of the present application for a method of obtaining external shape difference in aircraft component assembly.

[0071] Figure descriptions: 1. Skeleton part a; 2. Skeleton part b; 3. Skeleton part c; 4. Aircraft skin; 5. Aircraft component; 6. Typical skeleton part; 7. Gasket d; 8. Gasket e; 9. Outer skin; 10. Skeleton part f; 11. Skeleton part g.

[0072] The realization of the purpose, functional features and advantages of this application will be further explained in conjunction with the embodiments and with reference to the accompanying drawings. Detailed Implementation

[0073] It should be understood that the specific embodiments described herein are merely illustrative of this application and are not intended to limit this application.

[0074] The main solution of this application embodiment is: a method, apparatus, device, and storage medium for obtaining the external shape error of an aircraft component assembly. This involves acquiring process assembly information of the analysis area of ​​the external shape error of a target aircraft component; wherein the process assembly information includes assembly tolerances; obtaining the assembly relationship chain of the analysis area based on the process assembly information; wherein the assembly relationship chain is the aircraft component assembly relationship from the analysis datum to the analysis area; constructing an error calculation model based on the assembly relationship chain; obtaining a prediction index of the external shape error of the target aircraft component based on the error calculation model; and obtaining the assembly external shape error value of the target aircraft component based on the prediction index.

[0075] In existing technologies, the accuracy of an aircraft's shape is crucial to its aerodynamic characteristics, significantly impacting its flight performance and quality. Both aircraft designers and manufacturers strive for an actual aircraft shape that closely approximates the theoretical shape. However, due to unavoidable factors such as manufacturing errors, assembly errors, and deformation caused by the aircraft's own weight, the actual shape of the manufactured aircraft will inevitably deviate from the theoretical shape.

[0076] In aircraft manufacturing, process engineers pre-design and allocate manufacturing tolerances for parts and assembly tolerances for components. They then inspect the aircraft's external shape using inspection fixtures or digital measurement methods (such as laser trackers), adjust tolerances based on the inspection results, and use these tolerances to adjust the aircraft's external shape to ensure the accuracy of the components' shapes. However, current tolerance adjustments lack theoretical basis and rely primarily on the experience of process engineers, resulting in relatively poor accuracy. Furthermore, there is a conflict between the quality and efficiency of the adjustments: the process of adjusting and controlling the accuracy of the aircraft's external shape cannot be completed in one step; it requires repeated measurements and inspections over several runs. From an efficiency standpoint, fewer repetitions are better; however, from a data quality standpoint, the more measurements used to support the accuracy of the external shape data, the better.

[0077] To address this, this application provides a solution that, based on the actual assembly relationships of an aircraft and considering both manufacturing and assembly errors, employs a three-dimensional tolerance modeling method for prediction. This avoids relying on experience for judgment and improves the accuracy of adjustments. Furthermore, it uses simulation calculations to predict the shape of aircraft components, eliminating the need for tooling that requires measuring and judging the actual shape of the aircraft, resulting in lower costs and higher efficiency. Moreover, it can predict aircraft shape errors based on existing manufacturing tolerance design error models, or it can design process tolerances for manufacturing process tolerances based on the target aircraft shape error requirements, avoiding the repeated trial and error and adjustment process of traditional methods, thus balancing quality and efficiency.

[0078] Reference Figure 1 , Figure 1 This is a schematic diagram of the computer device structure of the hardware operating environment involved in the embodiments of this application.

[0079] like Figure 1As shown, the computer device may include: a processor 1001, such as a central processing unit (CPU), a communication bus 1002, a user interface 1003, a network interface 1004, and a memory 1005. The communication bus 1002 is used to enable communication between these components. The user interface 1003 may include a display screen and an input unit such as a keyboard; optionally, the user interface 1003 may also include a standard wired interface or a wireless interface. The network interface 1004 may optionally include a standard wired interface or a wireless interface (such as a Wi-Fi interface). The memory 1005 may be a high-speed random access memory (RAM) or a stable non-volatile memory (NVM), such as a disk drive. The memory 1005 may also optionally be a storage device independent of the aforementioned processor 1001.

[0080] Those skilled in the art will understand that Figure 1 The structure shown does not constitute a limitation on the computer device and may include more or fewer components than shown, or combine certain components, or have different component arrangements.

[0081] like Figure 1 As shown, the memory 1005, which serves as a storage medium, may include an operating system, a data storage module, a network communication module, a user interface module, and electronic programs.

[0082] exist Figure 1 In the computer device shown, the network interface 1004 is mainly used for data communication with the network server; the user interface 1003 is mainly used for data interaction with the user; the processor 1001 and the memory 1005 in the computer device of the present invention can be set in the computer device, and the computer device calls the aircraft component assembly external appearance difference acquisition device stored in the memory 1005 through the processor 1001, and executes the aircraft component assembly external appearance difference acquisition method provided in the embodiment of this application.

[0083] Reference Figure 2 Based on the hardware device of the foregoing embodiments, embodiments of this application provide a method for obtaining the external appearance difference of aircraft component assembly, including:

[0084] S10: Obtain the process assembly information of the analysis area of ​​the target aircraft component shape error; wherein, the process assembly information includes assembly tolerances;

[0085] In the specific implementation process, the assembly relationship of aircraft components is sorted out, the shape error analysis area P of aircraft components is determined, and the process assembly information of this area is obtained.

[0086] This embodiment takes a certain component 5 of an aircraft as an example, and the structural configuration is as follows: Figure 10 As shown in Table 1, the relevant parameters of the process assembly information for the analysis area of ​​this component are listed.

[0087] Table 1

[0088]

[0089]

[0090] S20: Based on the process assembly information, obtain the assembly relationship chain of the analysis area; wherein, the assembly relationship chain is the aircraft component assembly relationship from the analysis datum to the analysis area;

[0091] In the specific implementation process, the assembly relationship chain of the analysis area is obtained based on the content of the process assembly information. The error analysis benchmark should be associated with the existing assembly benchmark of the aircraft component, and priority should be given to selecting from the existing assembly benchmark of the aircraft component, that is, the component assembly relationship from the analysis benchmark to the analysis area.

[0092] For details, please refer to Figure 11 In this embodiment, the datum C0 for analyzing the shape error of the aircraft component is first determined. The origin of C0 is (1470, 11270, -702) in the aircraft coordinate system, the X-axis direction vector is (1, 0, 0), the Y-axis direction vector is (0, 1, 0), and the Z-axis direction vector is (0, 0, 1). Figure 6 As shown, based on the design model or drawings, the assembly relationships of aircraft components are sorted out based on the process assembly information therein, and the assembly relationship chain C0-C1-C2-C3-C4-C5 from the analysis datum to the analysis area is obtained.

[0093] S30: Construct an error calculation model based on the assembly relationship chain;

[0094] In the specific implementation process, a calculation model for component shape error is constructed based on the assembly relationship chain. This error calculation model predicts aircraft shape error based on existing manufacturing tolerance information.

[0095] As an optional implementation, the step of constructing an error calculation model based on the assembly relationship chain includes:

[0096] S301: Based on the assembly relationship chain, obtain the first point vector and the first normal vector of any point in the analysis region in the reference coordinate system, as well as the second point vector and the second normal vector of that point in the analysis region coordinate system; wherein, the Z-axis of the analysis region coordinate system is in the same direction as the first normal vector;

[0097] In the specific implementation process, let the homogeneous coordinates of a point within the shape error analysis region P on the skeleton in the reference coordinate system C0 be... The homogeneous coordinates in the C5 coordinate system are: In the reference coordinate system C0 The homogeneous coordinates of the normal vector at point are The homogeneous coordinates in the Cn coordinate system are:

[0098] When creating the local coordinate system Cn, it is necessary to ensure that the Z-axis is aligned with the coordinate system. The arrow at the point They are moving in the same direction. (The reason for this alignment needs to be explained.)

[0099] Explanation: Local coordinate system C n Located on the outer surface of the last assembled skeleton part, subsequent external parts (such as skin) are assembled close to the outer surface of the skeleton part. Manufacturing errors in these external parts manifest as thickness deviations, specifically variations along the normal to the skeleton's outer surface. This method requires ensuring that the Z-axis is aligned with... The arrow at the point The purpose of using the same direction is to simplify the construction complexity of the skeleton shape error model, and to simplify the external assembly error from a matrix to a scalar accumulation along the normal direction of the skeleton shape surface.

[0100] S302: Based on the first point vector, the first normal vector, the second point vector, and the second normal vector, obtain the coordinate system transformation relationship under the theoretical digital model state;

[0101] In the specific implementation process, the coordinate system transformation relationship under the theoretical numerical model state can be obtained based on the coordinate information of the above points.

[0102] As an optional implementation, the step of obtaining the coordinate system transformation relationship under the theoretical numerical model state based on the first point vector, the first normal vector, the second point vector, and the second normal vector includes:

[0103] The coordinate system transformation relationship is obtained through the following formula:

[0104]

[0105]

[0106] refer to Figure 5 , Figure 6 As shown, the assembly consists of skeleton parts and external parts. Part 1 is the first skeleton part of the assembly. Part 2 is positioned and assembled onto Part 1, and so on. Part 3 is positioned and assembled onto the previous part. Part 4 is the shape of the external attachment part 3 of the skeleton. Let be the vector of the first point, and its coordinates are... Let the second point vector have coordinates of... Let be the first normal vector, and its coordinates are... Let be the first normal vector, and its coordinates are... The homogeneous matrix represents the transformation relationship from coordinate system C1 to coordinate system C0. The homogeneous matrix represents the transformation relationship from coordinate system C2 to coordinate system C1. The homogeneous matrix represents the transformation relationship from coordinate system C3 to coordinate system C2. C represents n coordinate system to C n-1 The homogeneous matrix representing the transformation relationship between coordinate systems.

[0107] In the specific implementation process, based on the coordinate information of the above points, the coordinate system transformation relationship under the theoretical numerical model state can be obtained:

[0108]

[0109]

[0110] in, The homogeneous matrix represents the transformation relationship from coordinate system C0 to coordinate system C1. The homogeneous matrix represents the transformation relationship from coordinate system C1 to coordinate system C2. The homogeneous matrix represents the transformation relationship from coordinate system C3 to coordinate system C2. The homogeneous matrix represents the transformation relationship from the C4 coordinate system to the C3 coordinate system. The homogeneous matrix represents the transformation relationship from the C5 coordinate system to the C4 coordinate system.

[0111]

[0112]

[0113]

[0114]

[0115]

[0116] When the local coordinate system C5 is created, its Z-axis is parallel to... The arrow at the point Same direction.

[0117] S303: Based on the coordinate system transformation relationship, construct the standard calculation model for the shape of aircraft components and the calculation model for the shape error of aircraft components respectively;

[0118] In the specific implementation process, the coordinate system is transformed according to the coordinate system transformation relationship under the theoretical numerical model state, and the standard calculation model of aircraft component shape and the error calculation model of aircraft component shape are constructed. These two models are used to calculate the error value of aircraft shape.

[0119] As an optional implementation, the step of constructing a standard calculation model for the aircraft component shape and a calculation model for the aircraft component shape error based on the coordinate system transformation relationship includes:

[0120] The standard calculation model for the shape of the aircraft component is constructed using the following relationship:

[0121]

[0122] Where, ε i The theoretical thickness of the outer skin 9, representing the frame of the aircraft component. The homogeneous matrix represents the transformation relationship from coordinate system C1 to coordinate system C0. The homogeneous matrix represents the transformation relationship from coordinate system C2 to coordinate system C1. The homogeneous matrix represents the transformation relationship from coordinate system C3 to coordinate system C2. C represents n coordinate system to C n-1 Homogeneous matrix relating coordinate system transformations;

[0123] The following relationship is used to construct a calculation model for the shape error of the aircraft component:

[0124]

[0125] Where, ε′ i The actual thickness of the outer skin of the aircraft component skeleton (9 m layers), Δδ n represent Error matrix The homogeneous matrix represents the transformation relationship from coordinate system C1 to coordinate system C0. The homogeneous matrix represents the transformation relationship from coordinate system C2 to coordinate system C1. The homogeneous matrix represents the transformation relationship from coordinate system C3 to coordinate system C2. C represents n coordinate system to C n-1 Homogeneous matrix relating coordinate system transformations;

[0126] In the specific implementation process, in this embodiment, the aircraft frame has 3 layers of outer skin 9, with theoretical thicknesses of ε1=1.375mm, ε2=1.0mm, and ε3=2.0mm respectively.

[0127]

[0128] N can be calculated based on the model. i =[780.105 -95.032 812.129 1] T .

[0129] Construct a computational model of the aircraft component's shape that includes manufacturing and assembly errors:

[0130]

[0131] Where ε′1 represents the actual thickness of gasket d7, ε′2 represents the actual thickness of gasket e8, and ε′3 represents the actual thickness of outer skin 9, with the following values:

[0132]

[0133] Specifically, in the calculation model for the standard shape of aircraft components and the calculation model for the shape error of aircraft components, Δδ n represent The error matrix of the assembly-related matrix is ​​constructed similarly:

[0134]

[0135] refer to Figure 7 , where u n v n w n α n β n γ n The error matrix parameters form the error matrix parameter vector X. i =(u i v i w i α i β i γ i ), i = 1, 2, ..., n, the specific value of this vector depends on the feature form, assembly positioning form and tolerance value t. i .

[0136] Error matrix parameter vector Xi With tolerance t i The relationship between them is as follows:

[0137] When the shape of an aircraft component is a planar form and position feature, refer to Figure 8 , t i L1 and L2 represent the long and short side values ​​of the planar feature region, respectively, and the error matrix parameter vector is:

[0138]

[0139]

[0140] When the shape of an aircraft component is cylindrical, t i H represents the tolerance of the planar feature, and H represents the length of the cylindrical feature. The error matrix parameter vector is:

[0141]

[0142]

[0143] In this embodiment, Δδ1 to Δδ5 represent assembly relationship matrices, respectively. The error matrix is:

[0144]

[0145] Among them, u i v i w i α i β i γ i The parameter is the error matrix parameter, corresponding to the assembly or geometric tolerance t. i C1 and C2 form a planar uteronomy, C3 and C4 form a planar uteronomy, and the local feature where C5 is located is approximately a small plane. The specific composition of the error matrix parameter vector is as follows:

[0146]

[0147]

[0148]

[0149]

[0150]

[0151] S304: Construct the error calculation model based on the standard calculation model of the aircraft component shape and the error calculation model of the aircraft component shape.

[0152] In the specific implementation process, an error calculation model for point Ni of the aircraft component shape is constructed based on the standard calculation model of the aircraft component shape and the error calculation model of the aircraft component shape. This error calculation model is used to identify the error at this point in the future.

[0153] As an optional implementation, the step of constructing the error calculation model based on the aircraft component shape standard calculation model and the aircraft component shape error calculation model includes:

[0154] The error calculation model is constructed using the following relationship:

[0155]

[0156] Among them, M i For the calculation model of aircraft component shape error, N i A standard calculation model for the shape of aircraft components.

[0157] In the specific implementation process, the calculation model for the shape error at point Ni of the aircraft component is constructed as follows: in, Let M be the first normal vector. i For the calculation model of aircraft component shape error, N i A standard calculation model for the shape of aircraft components.

[0158] S40: Based on the error calculation model, obtain the predicted index of the shape error of the target aircraft component;

[0159] In practical implementation, the prediction indicators refer to the mean and standard deviation of the data in the error calculation dataset, which are used for subsequent judgment of the pass rate, etc. This prediction method is applicable to scenarios such as manufacturing error prediction or process tolerance design, and has good promotional value.

[0160] As an optional implementation, the step of obtaining the predicted index of the shape error of the target aircraft component based on the error calculation model includes:

[0161] S401: Obtain the error matrix based on the assembly tolerances;

[0162] In the specific implementation process, according to the assembly tolerance t i For the error matrix parameter vector {X i Data is sampled from {i = 1, 2, ..., n}, and a transformation error matrix {Δδ} is constructed based on the sampled values. i For the sequence number i = 1, 2, ..., n, the sampling method is to sample the error parameter according to a normal distribution, i.e. In this embodiment, the error matrix parameter vector is {X} iThe tolerance range of the error parameter is 6 times the standard deviation of the normal distribution, i.e.:

[0163]

[0164]

[0165]

[0166]

[0167]

[0168] according to The transformation yields the error matrix {Δδ} i , i = 1, 2, ..., 5}.

[0169] S402: Obtain the actual thickness of the outer skin of the aircraft component frame;

[0170] In a specific implementation, the actual thickness value ε′ of the outer skin 9 of the aircraft component error is used. i Data sampling was performed using a normal distribution for the actual thickness parameters of the outer skin 9. The sampling standard deviation is 1 / 6 of the tolerance range of the error parameter, that is...

[0171] Specifically, in this embodiment, data sampling is performed according to the above method, and the result is as follows:

[0172] ε′1~N(1.375, 0.11 / 6)

[0173] ε′2~N(1, 0.08 / 6)

[0174] ε′3~N(2, 0.16 / 6)

[0175] S403: Based on the error matrix and the actual thickness of the skin, perform simulation calculations on the aircraft component shape error calculation model to obtain the aircraft component shape error simulation dataset;

[0176] In the specific implementation process, the Monte Carlo method is adopted. Based on the error matrix obtained in steps S401 and S402, simulation calculations are performed on the aircraft component shape error calculation model obtained in step S303 to obtain the aircraft component shape simulation dataset {M}. i Let i = 1, 2, ..., N, where N is the number of sampling calculations in the Monte Carlo simulation. The Monte Carlo method is used for simulation calculations, and the results are statistically significant and robust. In this embodiment, the number of sampling calculations in the Monte Carlo simulation is 10,000.

[0177] S404: Obtain the error calculation dataset based on the aircraft component shape error simulation dataset and the error calculation model;

[0178] In the specific implementation process, based on the aircraft component shape simulation dataset M i Sum of error calculation model E i The error calculation dataset is obtained. In this embodiment, the set of error calculation values ​​at point Ni on the aircraft component is {E}. i |E i ∈E,i=1,2,...,10000}.

[0179] S405: Obtain the prediction index based on the error calculation dataset; wherein the prediction index is the mean and standard deviation of the data in the error calculation dataset.

[0180] In practical implementation, the predictive metrics refer to the mean and standard deviation of the data in the error calculation dataset. Based on the dataset {E} i} Calculate the mean μ, standard deviation σ, and maximum value E. max Minimum value E min , median E min Thus, a predictive index for the shape error at point Ni on the aircraft component was obtained.

[0181]

[0182]

[0183] In this embodiment, the relevant data for predicting the shape error at point Ni on the upper part of a certain component 5 of the aircraft to be tested are as follows: average shape error μ = -0.073, standard deviation of shape error σ = 0.226, and maximum shape error E. max =0.923, Minimum value of shape error E min = -0.949, median of shape error E mid = -0.070.

[0184] S50: Based on the predicted index, obtain the assembly appearance difference of the target aircraft component.

[0185] In the specific implementation process, the manufacturing process tolerance pass rate is calculated based on the predicted index of component shape error, and the assembly shape error value of the target aircraft component is obtained based on the manufacturing process tolerance pass rate that meets the manufacturing requirements.

[0186] As an optional implementation, the step of obtaining the assembly appearance difference of the target aircraft component based on the predicted index includes:

[0187] Determine whether the manufacturing tolerance pass rate of the assembly tolerance is greater than or equal to meeting the preset process requirements; wherein, the manufacturing tolerance pass rate is obtained based on the prediction index;

[0188] If so, the assembly tolerance is the difference in the external appearance of the assembly;

[0189] If not, the assembly tolerance corresponding to the error parameter with the largest contribution in the error matrix is ​​selected for adjustment to obtain the assembly external tolerance value.

[0190] In the specific implementation process, the manufacturing process tolerance pass rate C is calculated based on the statistical indicators of the calculated component shape error values. pk Determine whether the manufacturing tolerance pass rate is greater than or equal to the preset process requirement U. The value of U is determined by the design or process requirements, and usually U≥1: when C pk When C ≥ U, it is determined that the tolerance parameters of the current manufacturing process can meet the external shape tolerance requirements of the aircraft parts, and the assembly tolerance at this time is the assembly external shape tolerance value; when C pk When <U, it is determined that the current manufacturing process tolerance parameter cannot meet the aircraft component shape tolerance requirements, and the tolerance parameter needs to be adjusted. That is, the assembly tolerance corresponding to the error parameter with the largest contribution in the error matrix is ​​selected for adjustment in order to obtain the assembly shape tolerance value.

[0191] Specifically, in this embodiment, C pk =0.629, U value is 1, C pk <U indicates that the tolerance parameters of the current manufacturing process need to be adjusted.

[0192] As an optional implementation, the manufacturing tolerance pass rate can be obtained through the following relationship:

[0193]

[0194] Among them, T max For the upper limit of the shape error of aircraft components, T min Here, μ and σ represent the lower limit of the shape error of aircraft components, and μ is the prediction index. σ is the standard deviation.

[0195] In the specific implementation process, the manufacturing tolerance pass rate in this embodiment is:

[0196] As an optional implementation, the contribution level can be obtained through the following relationship:

[0197]

[0198] Among them, ti For assembly tolerances, σ is the standard deviation.

[0199] In the specific implementation process, the contribution of each error parameter is calculated based on statistical indicators. The calculation method is as follows: The calculation results of the contribution are shown in Table 2:

[0200] Table 2

[0201]

[0202] At this point, the step of selecting the assembly tolerance corresponding to the error parameter with the largest contribution in the error matrix for adjustment to obtain the assembly external tolerance value includes:

[0203] Modify the tolerance zone position and tolerance zone width of the assembly tolerance corresponding to the error parameter with the largest contribution to obtain the assembly external tolerance value.

[0204] In the specific implementation process, if the manufacturing tolerance pass rate of the assembly tolerance is less than the preset process requirements, the assembly tolerance corresponding to the error parameter with the largest contribution in the error matrix is ​​selected for adjustment, and the above steps S40 and S50 are repeated until the obtained C is obtained. pk The value meets the U-value requirement. The specific adjustment method is as follows:

[0205] 1. Select the assembly tolerance t corresponding to the error parameter with the largest contribution. i ;

[0206] 2. Adjust and modify the tolerance zone position of the assembly tolerance, and re-obtain the error parameter X. i By analyzing the distribution pattern of ε′1, we can obtain the predictive index: the average value. Standard deviation Repeat steps S40 and S50 above to calculate the symmetry index E of the shape error. sym until satisfied:

[0207]

[0208] This indicator is used to adjust the tolerance zone position of assembly tolerances.

[0209] 3. Adjust and modify the tolerance zone width of the assembly tolerances so that the recalculated C... pk The value meets the U-value requirement.

[0210] During the adjustment process, assembly tolerances can be adjusted by adding or subtracting increments, with an adjustment increment of 0.001 mm, but the following condition must be met: assembly tolerance t i The standard deviation of the corresponding generated error parameters

[0211] Generally, any manufacturing tolerance has upper and lower limits. The smaller the range of these limits, the narrower the tolerance, which means higher manufacturing precision requirements and correspondingly higher manufacturing costs. The aforementioned restrictions in this application are intended to control the tolerance range of a particular assembly from becoming too small, thereby limiting overly stringent manufacturing precision control and minimizing manufacturing costs.

[0212] In this embodiment, based on the contribution calculation results, the assembly tolerances t1, t3, and t5 are adjusted, and the above steps S40 and S50 are repeated to recalculate the appearance error statistical index and re-determine the manufacturing process tolerance pass rate C. pk Whether the requirement of ≥U is met, the error parameters after the above steps of adjustment and optimization are shown in Table 3:

[0213] Table 3

[0214]

[0215] The relevant data for the shape error prediction index at this time are: average shape error μ = 0.001, standard deviation of shape error σ = 0.166, and maximum shape error E. max =0.634, Minimum value of shape error E min = -0.649, median of shape error E mid = 0.002. At this point, the symmetry index E of the shape error is... sym =0.995, Manufacturing process tolerance pass rate C pk =1.005, satisfying C pk The requirement is ≥U.

[0216] Adjusted assembly tolerance t i This refers to the final assembly external shape tolerance value {S} that meets the shape error requirements of aircraft components. i}

[0217] It should be understood that the above are merely illustrative examples and do not constitute any limitation on the technical solution of this application. Those skilled in the art can make settings as needed in practical applications, and no restrictions are imposed here.

[0218] As can be seen from the above description, this embodiment is based on the actual assembly relationship of the aircraft, considers the manufacturing and assembly errors of parts, and uses a three-dimensional tolerance modeling method for prediction, avoiding the situation of relying on experience for judgment and improving the accuracy of adjustment. Furthermore, it uses simulation calculation method to predict the shape of aircraft components, eliminating the need for tooling that requires measuring and judging the actual shape of the aircraft, which is low-cost and highly efficient. In addition, it can predict the shape error of the aircraft based on the existing manufacturing tolerance design error model, or it can design the process tolerance of the manufacturing process based on the target aircraft shape error requirements, avoiding the process of repeated trial and error and adjustment of traditional methods, and taking into account both quality and efficiency.

[0219] Reference Figure 3 Based on the same inventive concept, embodiments of this application also provide an aircraft component assembly external appearance difference acquisition device, comprising:

[0220] The process assembly information acquisition module is used to acquire process assembly information of the analysis area of ​​the shape error of the target aircraft component; wherein, the process assembly information includes assembly tolerances.

[0221] The assembly relationship chain acquisition module is used to acquire the assembly relationship chain of the analysis area based on the process assembly information; wherein, the assembly relationship chain is the aircraft component assembly relationship from the analysis datum to the analysis area;

[0222] An error calculation model construction module is used to construct an error calculation model based on the assembly relationship chain;

[0223] The prediction index acquisition module is used to acquire the prediction index of the shape error of the target aircraft component according to the error calculation model.

[0224] The tolerance value acquisition module is used to acquire the assembly external tolerance value of the target aircraft component based on the prediction index.

[0225] It should be noted that each module in the aircraft component assembly external appearance difference acquisition device in this embodiment corresponds one-to-one with each step in the aircraft component assembly external appearance difference acquisition method in the aforementioned embodiment. Therefore, the specific implementation of this embodiment can refer to the implementation of the aforementioned aircraft component assembly external appearance difference acquisition method, and will not be repeated here.

[0226] Furthermore, in one embodiment, the present application also provides a computer device, the device including a processor, a memory, and a computer program stored in the memory, the computer program being executed by the processor to implement the steps of the methods in the foregoing embodiments.

[0227] Furthermore, in one embodiment, the present application also provides a computer storage medium storing a computer program, which, when executed by a processor, implements the steps of the methods described in the foregoing embodiments.

[0228] In some embodiments, the computer-readable storage medium may be a memory such as FRAM, ROM, PROM, EPROM, EEPROM, flash memory, magnetic surface memory, optical disk, or CD-ROM; or it may be a device including one or any combination of the above-mentioned memories. The computer may be a variety of computing devices, including smart terminals and servers.

[0229] In some embodiments, executable instructions may take the form of a program, software, software module, script, or code, written in any form of programming language (including compiled or interpreted languages, or declarative or procedural languages), and may be deployed in any form, including as a standalone program or as a module, component, subroutine, or other unit suitable for use in a computing environment.

[0230] As an example, executable instructions may, but do not necessarily, correspond to files in a file system. They may be stored as part of a file that holds other programs or data, for example, in one or more scripts in a Hyper Text Markup Language (HTML) document, in a single file dedicated to the program in question, or in multiple collaborating files (e.g., a file that stores one or more modules, subroutines, or code sections).

[0231] As an example, executable instructions can be deployed to execute on a single computing device, or on multiple computing devices located in one location, or on multiple computing devices distributed across multiple locations and interconnected via a communication network.

[0232] It should be noted that, in this document, the terms "comprising," "including," or any other variations thereof are intended to cover non-exclusive inclusion, such that a process, method, article, or system that comprises a list of elements includes not only those elements but also other elements not expressly listed, or elements inherent to such a process, method, article, or system. Unless otherwise specified, an element defined by the phrase "comprising one..." does not exclude the presence of other identical elements in the process, method, article, or system that includes that element.

[0233] The sequence numbers of the embodiments in this application are for descriptive purposes only and do not represent the superiority or inferiority of the embodiments.

[0234] Through the above description of the embodiments, those skilled in the art can clearly understand that the methods of the above embodiments can be implemented by means of software plus necessary general-purpose hardware platforms. Of course, they can also be implemented by hardware, but in many cases the former is a better implementation method. Based on this understanding, the technical solution of this application, in essence, or the part that contributes to the prior art, can be embodied in the form of a software product. This computer software product is stored in a storage medium (such as read-only memory / random access memory, magnetic disk, optical disk) and includes several instructions to cause a multimedia terminal device (which may be a mobile phone, computer, television receiver, or network device, etc.) to execute the methods described in the various embodiments of this application.

[0235] The above are merely preferred embodiments of this application and do not limit the patent scope of this application. Any equivalent structural or procedural transformations made using the content of this application's specification and drawings, or direct or indirect applications in other related technical fields, are similarly included within the patent protection scope of this application.

Claims

1. A method for obtaining the external shape difference of an aircraft component assembly, characterized in that, Includes the following steps: Obtain process assembly information for the analysis area of ​​the shape error of the target aircraft component; wherein, the process assembly information includes assembly tolerances; Based on the process assembly information, the assembly relationship chain of the analysis area is obtained; wherein, the assembly relationship chain is the aircraft component assembly relationship from the analysis datum to the analysis area; Based on the assembly relationship chain, obtain the first point vector and the first normal vector of any point in the analysis region in the reference coordinate system, as well as the second point vector and the second normal vector of that point in the analysis region coordinate system; wherein, the Z-axis of the analysis region coordinate system is in the same direction as the first normal vector; based on the first point vector, the first normal vector, the second point vector, and the second normal vector, obtain the coordinate system transformation relationship under the theoretical numerical model state; based on the coordinate system transformation relationship, construct the standard calculation model for the aircraft component shape and the error calculation model for the aircraft component shape respectively; based on the standard calculation model for the aircraft component shape and the error calculation model for the aircraft component shape, construct the error calculation model; Based on the error calculation model, the predicted index of the shape error of the target aircraft component is obtained; Based on the predicted indicators, the assembly appearance difference of the target aircraft component is obtained; The steps of constructing the standard calculation model of aircraft component shape and the calculation model of aircraft component shape error based on the coordinate system transformation relationship include: The standard calculation model for the shape of the aircraft component is constructed using the following relationship: in, The theoretical thickness representing the outer skin of an aircraft component's frame; express coordinate system to Homogeneous matrix representing coordinate system transformations. express coordinate system to Homogeneous matrix representing coordinate system transformations. express coordinate system to Homogeneous matrix representing coordinate system transformations. express coordinate system to Homogeneous matrix relating coordinate system transformations; The following relationship is used to construct a calculation model for the shape error of the aircraft component: in, This represents the actual thickness of the skin covering the outer frame of an aircraft component. represent The error matrix, express coordinate system to Homogeneous matrix representing coordinate system transformations. express coordinate system to Homogeneous matrix representing coordinate system transformations. express coordinate system to Homogeneous matrix representing coordinate system transformations. express coordinate system to Homogeneous matrix relating coordinate system transformations; The step of constructing the error calculation model based on the aircraft component shape standard calculation model and the aircraft component shape error calculation model includes: The error calculation model is constructed using the following relationship: in, A model for calculating the shape error of aircraft components. A standard calculation model for the shape of aircraft components.

2. The method for obtaining the external appearance difference of aircraft component assembly as described in claim 1, characterized in that, The step of obtaining the coordinate system transformation relationship under the theoretical numerical model state based on the first point vector, the first normal vector, the second point vector, and the second normal vector includes: The coordinate system transformation relationship is obtained through the following formula: in, Let be the vector of the first point, and its coordinates are... ; Let the second point vector have coordinates of... ; Let be the first normal vector, and its coordinates are... ; Let be the second normal vector, and its coordinates be... ; express coordinate system to Homogeneous matrix representing coordinate system transformations. express coordinate system to Homogeneous matrix representing coordinate system transformations. express coordinate system to Homogeneous matrix representing coordinate system transformations. express coordinate system to The homogeneous matrix representing the transformation relationship between coordinate systems.

3. The method for obtaining the external appearance difference of aircraft component assembly as described in claim 1, characterized in that, The step of obtaining the predicted index of the shape error of the target aircraft component based on the error calculation model includes: Based on the assembly tolerances, obtain the error matrix; Obtain the actual thickness of the outer skin of the aircraft component frame; Based on the error matrix and the actual thickness of the skin, the aircraft component shape error calculation model is simulated to obtain the aircraft component shape error simulation dataset. Based on the aircraft component shape error simulation dataset and the error calculation model, an error calculation dataset is obtained; The prediction index is obtained based on the error calculation dataset; wherein the prediction index is the mean and standard deviation of the data in the error calculation dataset.

4. The method for obtaining the external appearance difference of aircraft component assembly as described in claim 1, characterized in that, The step of obtaining the assembly appearance difference value of the target aircraft component based on the predicted index includes: Determine whether the manufacturing tolerance pass rate of the assembly tolerance is greater than or equal to meeting the preset process requirements; wherein, the manufacturing tolerance pass rate is obtained based on the prediction index; If so, the assembly tolerance is the difference in the external appearance of the assembly; If not, the assembly tolerance corresponding to the error parameter with the largest contribution in the error matrix is ​​selected for adjustment to obtain the assembly external tolerance value.

5. The method for obtaining the external appearance difference of aircraft component assembly as described in claim 4, characterized in that, The manufacturing tolerance pass rate is obtained through the following relationship: in, This refers to the upper limit of the dimensional error of aircraft components. This refers to the lower limit of the shape error of aircraft components. and As a predictive indicator, This is the average value. , Standard deviation ; Here is the error calculation model, and N is the number of sampling calculations in the Monte Carlo simulation.

6. The method for obtaining the external appearance difference of aircraft component assembly as described in claim 5, characterized in that, The contribution level is obtained through the following relationship: Among them, among them, For assembly tolerances, For assembly tolerances The standard deviation of the corresponding generated error parameter, Standard deviation ; If not, the step of selecting the assembly tolerance corresponding to the error parameter with the largest contribution in the error matrix for adjustment to obtain the assembly external tolerance value includes: Modify the tolerance zone position and tolerance zone width of the assembly tolerance corresponding to the error parameter with the largest contribution to obtain the assembly external tolerance value.

7. A device for obtaining external shape difference of aircraft component assembly, characterized in that, For implementing the method as described in claim 1, the apparatus comprises: The process assembly information acquisition module is used to acquire process assembly information of the analysis area of ​​the shape error of the target aircraft component; wherein, the process assembly information includes assembly tolerances. The assembly relationship chain acquisition module is used to acquire the assembly relationship chain of the analysis area based on the process assembly information; wherein, the assembly relationship chain is the aircraft component assembly relationship from the analysis datum to the analysis area; An error calculation model construction module is used to construct an error calculation model based on the assembly relationship chain; The prediction index acquisition module is used to acquire the prediction index of the shape error of the target aircraft component according to the error calculation model. The tolerance value acquisition module is used to acquire the assembly external tolerance value of the target aircraft component based on the prediction index.

8. A computer device, characterized in that, The computer device includes a memory and a processor, wherein the memory stores a computer program and the processor executes the computer program to implement the method as described in any one of claims 1-6.

9. A computer-readable storage medium, characterized in that, The computer-readable storage medium stores a computer program, and the processor executes the computer program to implement the method as described in any one of claims 1-6.

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