CN117349955A - Method, device, equipment and storage medium for acquiring assembly shape tolerance of aircraft component - Google Patents

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

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CN117349955A
CN117349955A CN202311164503.8A CN202311164503A CN117349955A CN 117349955 A CN117349955 A CN 117349955A CN 202311164503 A CN202311164503 A CN 202311164503A CN 117349955 A CN117349955 A CN 117349955A
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assembly
error
aircraft component
tolerance
coordinate system
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陈雪梅
陈清良
勾江洋
骆金威
潘雨
刘元吉
李栎森
叶翔宇
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Chengdu Aircraft Industrial Group Co Ltd
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Chengdu Aircraft Industrial Group Co Ltd
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    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64F—GROUND OR AIRCRAFT-CARRIER-DECK INSTALLATIONS SPECIALLY ADAPTED FOR USE IN CONNECTION WITH AIRCRAFT; DESIGNING, MANUFACTURING, ASSEMBLING, CLEANING, MAINTAINING OR REPAIRING AIRCRAFT, NOT OTHERWISE PROVIDED FOR; HANDLING, TRANSPORTING, TESTING OR INSPECTING AIRCRAFT COMPONENTS, NOT OTHERWISE PROVIDED FOR
    • B64F5/00—Designing, manufacturing, assembling, cleaning, maintaining or repairing aircraft, not otherwise provided for; Handling, transporting, testing or inspecting aircraft components, not otherwise provided for
    • B64F5/10—Manufacturing or assembling aircraft, e.g. jigs therefor
    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F30/00—Computer-aided design [CAD]
    • G06F30/10—Geometric CAD
    • G06F30/15—Vehicle, aircraft or watercraft design
    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F30/00—Computer-aided design [CAD]
    • G06F30/10—Geometric CAD
    • G06F30/17—Mechanical parametric or variational design
    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F30/00—Computer-aided design [CAD]
    • G06F30/20—Design optimisation, verification or simulation

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Abstract

The application discloses a method, a device, equipment and a storage medium for acquiring the assembly appearance tolerance of an aircraft component, which are based on the real assembly relation of the aircraft, consider the manufacturing error and the assembly error of parts, predict by adopting a three-dimensional tolerance modeling method, avoid the condition of judging by relying on experience and improve the accuracy of adjustment; moreover, the appearance of the aircraft component is predicted by adopting a simulation calculation method, so that a tool for measuring and judging the real appearance of the aircraft is eliminated, the cost is low, and the efficiency is high; furthermore, the aircraft appearance error can be predicted based on the existing manufacturing tolerance design error model, and the manufacturing process tolerance can be designed based on the target aircraft appearance error requirement, so that the processes of repeated trial and error and adjustment of the traditional method are avoided, and the quality and efficiency are both considered.

Description

Method, device, equipment and storage medium for acquiring assembly shape tolerance of aircraft component
Technical Field
The present disclosure relates to the field of aircraft assembly, and in particular, to a method, an apparatus, a device, and a storage medium for obtaining an assembly profile tolerance of an aircraft component.
Background
The accuracy of the appearance of the aircraft relates to the aerodynamic characteristics of the aircraft, and has great influence on the flight performance and the flight quality of the aircraft. The actual appearance of the aircraft can be close to the theoretical appearance, which is expected by both aircraft designers and aircraft manufacturers, however, due to unavoidable factors such as part manufacturing errors, assembly errors, dead weight deformation of the aircraft and the like, the actually produced aircraft appearance is necessarily deviated from the theoretical appearance.
In the aircraft manufacturing process, the process personnel perform design allocation on the manufacturing tolerance of the parts and the assembly tolerance of the parts in advance, the appearance of the aircraft is checked by using an inspection tool or a digital measurement means (such as a laser tracker and the like), the tolerance is adjusted according to the appearance checking result, and the appearance of the aircraft is adjusted by the tolerance, so that the accuracy of the appearance of the parts of the aircraft is ensured. However, the adjustment of the tolerance at present lacks theoretical basis, and is mainly carried out by relying on the working experience of the process staff, so that the accuracy is poor.
Disclosure of Invention
The main purpose of the application is to provide a method, a device, equipment and a storage medium for obtaining the assembly appearance tolerance of an aircraft component, and aims to solve the technical problem that the accuracy of the existing tolerance adjustment method is poor.
To achieve the above object, the present application provides a method for obtaining an assembly profile tolerance of an aircraft component, including:
acquiring process assembly information of an analysis area of the appearance error of the target aircraft component; wherein, the process assembly information comprises assembly tolerance;
acquiring an assembly relation chain of the analysis area according to the process assembly information; wherein the assembly relationship chain is an aircraft component assembly relationship from an analysis reference to the analysis region;
constructing an error calculation model according to the assembly relation chain;
obtaining a prediction index of the appearance error of the target aircraft component according to the error calculation model;
and acquiring the assembly appearance tolerance value of the target aircraft component according to the prediction index.
Optionally, the step of constructing an error calculation model according to the assembly relation chain includes:
according to the assembly relation chain, a first point vector and a first normal vector of any point in the analysis area under a reference coordinate system, and a second point vector and a second normal vector of the point under the analysis area coordinate system are obtained; wherein the Z axis of the analysis region coordinate system is in the same direction as the first normal vector;
Acquiring a coordinate system conversion relation in a theoretical digital-analog state according to the first point vector, the first normal vector, the second point vector and the second normal vector;
respectively constructing an aircraft component appearance standard calculation model and an aircraft component appearance error calculation model according to the coordinate system conversion relation;
and constructing the error calculation model according to the aircraft component appearance standard calculation model and the aircraft component appearance error calculation model.
Optionally, the step of obtaining the coordinate system conversion relationship in the theoretical digital-analog state according to the first point vector, the first normal vector, the second point vector and the second normal vector includes:
the coordinate system conversion relation is obtained through the following relation:
wherein,is a first point vector whose coordinate value is +.> Is a second point vector whose coordinate value is +.> Is a first normal vector with coordinate values of Is a first normal vector whose coordinate value is +.> Represent C 1 Coordinate system to C 0 Conversion relation homogeneous matrix of coordinate system, +.>Represent C 2 Coordinate system to C 1 Conversion relation homogeneous matrix of coordinate system, +.>Represent C 3 Coordinate system to C 2 Conversion relation homogeneous matrix of coordinate system, +. >Represent C n Coordinate system to C n-1 And (5) a homogeneous matrix of the conversion relation of the coordinate system.
Optionally, the step of respectively constructing an aircraft component shape standard calculation model and an aircraft component shape error calculation model according to the coordinate system conversion relation includes:
constructing the aircraft component appearance standard calculation model by the following relation:
wherein ε i Represents the theoretical thickness of the outer skin of the skeleton of the aircraft component,Represent C 1 Coordinate system to C 0 Conversion relation homogeneous matrix of coordinate system, +.>Represent C 2 Coordinate system to C 1 Conversion relation homogeneous matrix of coordinate system, +.>Represent C 3 Coordinate system to C 2 Conversion relation homogeneous matrix of coordinate system, +.>Represent C n Coordinate system to C n-1 A homogeneous matrix of the conversion relation of the coordinate system;
the aircraft component shape error calculation model is constructed by the following relation:
wherein ε' i Representing the actual thickness of the skin (m layers) outside the skeleton of an aircraft component, delta n Representative ofError matrix of (2),Represent C 1 Coordinate system to C 0 Conversion relation homogeneous matrix of coordinate system, +.>Represent C 2 Coordinate system to C 1 Conversion relation homogeneous matrix of coordinate system, +.>Represent C 3 Coordinate system to C 2 Conversion relation homogeneous matrix of coordinate system, +.>Represent C n Coordinate system to C n-1 A homogeneous matrix of the conversion relation of the coordinate system;
The step of constructing the error calculation model according to the aircraft component appearance standard calculation model and the aircraft component appearance error calculation model comprises the following steps:
the error calculation model is constructed by the following relation:
wherein M is i Model N for calculating shape error of aircraft component i A model is calculated for the aircraft part form standard.
Optionally, the step of obtaining a prediction index of the shape error of the target aircraft component according to the error calculation model includes:
obtaining an error matrix according to the assembly tolerance;
acquiring the actual thickness of the outer skin of the skeleton of the aircraft component;
based on the error matrix and the actual thickness of the skin, performing simulation calculation on an aircraft component shape error calculation model to obtain an aircraft component shape error simulation data set;
obtaining an error calculation data set according to the aircraft component shape error simulation data set and the error calculation model;
calculating a data set according to the error, and acquiring the prediction index; and calculating the average value and standard deviation of the data in the data set for the error by the prediction index.
Optionally, the step of obtaining the fitting appearance tolerance value of the target aircraft component according to the prediction index includes:
Judging whether the manufacturing tolerance qualification rate of the assembly tolerance is more than or equal to the preset process requirement; wherein, the manufacturing tolerance qualification rate is obtained according to the prediction index;
if yes, the assembly tolerance is the assembly external tolerance value;
if not, the assembly tolerance corresponding to the error parameter with the largest contribution degree in the error matrix is selected for adjustment, so that the assembly appearance tolerance value is obtained.
Optionally, the manufacturing tolerance yield is obtained by the following relation:
wherein T is max Is the upper limit value of the appearance error of the aircraft component, T min The lower limit value, mu and sigma of the appearance error of the aircraft component are used as prediction indexes, mu is an average value,sigma is standard deviation>
Optionally, the contribution is obtained by the following relation:
wherein t is i For assembly tolerances, σ is the standard deviation,
if not, selecting an assembly tolerance corresponding to an error parameter with the largest contribution degree in the error matrix for adjustment to obtain the assembly external tolerance value, wherein the step comprises the following steps:
and modifying the tolerance zone position and the tolerance zone width of the assembly tolerance corresponding to the error parameter with the largest contribution degree to obtain the assembly external tolerance value.
In addition, to achieve the above object, the present application further provides an aircraft component assembly profile tolerance obtaining apparatus, including:
The process assembly information acquisition module is used for acquiring process assembly information of an analysis area of the appearance error of the target aircraft component; wherein, the process assembly information comprises assembly tolerance;
the assembly relation chain acquisition module is used for acquiring the assembly relation chain of the analysis area according to the process assembly information; wherein the assembly relationship chain is an aircraft component assembly relationship from an analysis reference to the analysis region;
the error calculation model construction module is used for constructing an error calculation model according to the assembly relation chain;
the prediction index acquisition module is used for acquiring a prediction index of the appearance error of the target aircraft component according to the error calculation model;
and the tolerance value acquisition module is used for acquiring the assembly appearance tolerance value of the target aircraft component according to the prediction index.
In addition, to achieve the above object, the present application further provides a computer device, which includes a memory and a processor, where the memory stores a computer program, and the processor executes the computer program to implement the above method.
In addition, in order to achieve the above object, the present application further provides a computer readable storage medium, where a computer program is stored, and a processor executes the computer program to implement the above method.
The beneficial effects that this application can realize.
The method, the device, the equipment and the storage medium for acquiring the appearance tolerance of the assembly of the aircraft component are provided by the embodiment of the application, and the process assembly information of an analysis area of the appearance error of the target aircraft component is acquired; wherein, the process assembly information comprises assembly tolerance; acquiring an assembly relation chain of the analysis area according to the process assembly information; wherein the assembly relationship chain is an aircraft component assembly relationship from an analysis reference to the analysis region; constructing an error calculation model according to the assembly relation chain; obtaining a prediction index of the appearance error of the target aircraft component according to the error calculation model; and acquiring the assembly appearance tolerance value of the target aircraft component according to the prediction index. Based on the real assembly relation of the aircraft, the manufacturing error and the assembly error of the parts are considered, and the prediction is performed by adopting a three-dimensional tolerance modeling method, so that the situation of judging by experience is avoided, and the accuracy of adjustment is improved; moreover, the appearance of the aircraft component is predicted by adopting a simulation calculation method, so that a tool for measuring and judging the real appearance of the aircraft is eliminated, the cost is low, and the efficiency is high; furthermore, the aircraft appearance error can be predicted based on the existing manufacturing tolerance design error model, and the manufacturing process tolerance can be designed based on the target aircraft appearance error requirement, so that the processes of repeated trial and error and adjustment of the traditional method are avoided, and the quality and efficiency are both considered.
Drawings
FIG. 1 is a schematic diagram of a computer device in a hardware operating environment according to an embodiment of the present application;
FIG. 2 is a flow chart of a method for obtaining tolerance of an assembly profile of an aircraft component according to an embodiment of the present disclosure;
FIG. 3 is a functional block diagram of an aircraft component assembly profile tolerance acquisition device according to an embodiment of the present disclosure;
FIG. 4 is a schematic illustration of an aircraft component assembly profile tolerance acquisition method according to an embodiment of the present application;
FIG. 5 is a schematic exploded view of an aircraft component assembly profile tolerance acquisition method according to an embodiment of the present application;
FIG. 6 is an assembly relationship transfer chain diagram of an aircraft component assembly profile tolerance acquisition method provided in an embodiment of the present application;
fig. 7 is a schematic diagram of transformation of geometric feature position coordinates in different coordinate systems of an aircraft component assembly profile tolerance obtaining method according to an embodiment of the present application;
FIG. 8 is a schematic plan view of a method for obtaining tolerance of an assembly profile of an aircraft component according to an embodiment of the present application;
FIG. 9 is a schematic view of a cylindrical bit feature of an aircraft component assembly profile tolerance acquisition method provided in an embodiment of the present application;
FIG. 10 is an exploded view of an exemplary aircraft component assembly profile tolerance acquisition method according to an embodiment of the present application;
FIG. 11 is a schematic view of an exemplary component coordinate system of an aircraft component assembly profile tolerance acquisition method according to an embodiment of the present disclosure;
the description of the drawings is 1, a framework part a; 2. a skeleton part b; 3. a skeleton part c; 4. an aircraft skin; 5. a component of an aircraft; 6. typical skeleton parts; 7. a gasket d; 8. a gasket e; 9. an outer skin; 10. a framework part f; 11. and (3) a framework part g.
The realization, functional characteristics and advantages of the present application will be further described with reference to the embodiments, referring to the attached drawings.
Detailed Description
It should be understood that the specific embodiments described herein are for purposes of illustration only and are not intended to limit the present application.
The main solutions of the embodiments of the present application are: the method, the device, the equipment and the storage medium for acquiring the appearance tolerance of the assembly of the aircraft component are provided, and process assembly information of an analysis area of the appearance error of the target aircraft component is acquired; wherein, the process assembly information comprises assembly tolerance; acquiring an assembly relation chain of the analysis area according to the process assembly information; wherein the assembly relationship chain is an aircraft component assembly relationship from an analysis reference to the analysis region; constructing an error calculation model according to the assembly relation chain; obtaining a prediction index of the appearance error of the target aircraft component according to the error calculation model; and acquiring the assembly appearance tolerance value of the target aircraft component according to the prediction index.
In the prior art, the accuracy of the appearance of the aircraft relates to the aerodynamic characteristics of the aircraft, and the aircraft has great influence on the flight performance and the flight quality. The actual appearance of the aircraft can be close to the theoretical appearance, which is expected by both aircraft designers and aircraft manufacturers, however, due to unavoidable factors such as part manufacturing errors, assembly errors, dead weight deformation of the aircraft and the like, the actually produced aircraft appearance is necessarily deviated from the theoretical appearance.
In the aircraft manufacturing process, the process personnel perform design allocation on the manufacturing tolerance of the parts and the assembly tolerance of the parts in advance, the appearance of the aircraft is checked by using an inspection tool or a digital measurement means (such as a laser tracker and the like), the tolerance is adjusted according to the appearance checking result, and the appearance of the aircraft is adjusted by the tolerance, so that the accuracy of the appearance of the parts of the aircraft is ensured. However, the adjustment of the tolerance at present lacks theoretical basis, and is mainly carried out by relying on the working experience of the process staff, so that the accuracy is poor. And the quality and efficiency of the adjustment contradict: the aircraft appearance accuracy adjustment control process cannot be completed once, and a plurality of repeated measurement and inspection for the frames are needed, so that the fewer the repeated times, the better the efficiency; in terms of data quality, the more the number of measurement frames supporting the accuracy of the profile data, the better.
Therefore, the method and the device provide a solution, based on the real assembly relation of the aircraft, consider the manufacturing error and the assembly error of the parts, and predict by adopting a three-dimensional tolerance modeling method, so that the situation of judging by experience is avoided, and the accuracy of adjustment is improved; moreover, the appearance of the aircraft component is predicted by adopting a simulation calculation method, so that a tool for measuring and judging the real appearance of the aircraft is eliminated, the cost is low, and the efficiency is high; furthermore, the aircraft appearance error can be predicted based on the existing manufacturing tolerance design error model, and the manufacturing process tolerance can be designed based on the target aircraft appearance error requirement, so that the processes of repeated trial and error and adjustment of the traditional method are avoided, and the quality and efficiency are both considered.
Referring to fig. 1, fig. 1 is a schematic diagram of a computer device structure of a hardware running environment according to an embodiment of the present application.
As shown in fig. 1, the computer device may include: a processor 1001, such as a central processing unit (Central Processing Unit, CPU), a communication bus 1002, a user interface 1003, a network interface 1004, a memory 1005. Wherein the communication bus 1002 is used to enable connected communication between these components. The user interface 1003 may include a Display, an input unit such as a Keyboard (Keyboard), and the optional user interface 1003 may further include a standard wired interface, a wireless interface. The network interface 1004 may optionally include a standard wired interface, a WIreless interface (e.g., a WIreless-FIdelity (WI-FI) interface). The Memory 1005 may be a high-speed random access Memory (Random Access Memory, RAM) Memory or a stable nonvolatile Memory (NVM), such as a disk Memory. The memory 1005 may also optionally be a storage device separate from the processor 1001 described above.
Those skilled in the art will appreciate that the architecture shown in fig. 1 is not limiting of a computer device and may include more or fewer components than shown, or may combine certain components, or a different arrangement of components.
As shown in fig. 1, an operating system, a data storage module, a network communication module, a user interface module, and an electronic program may be included in the memory 1005 as one type of storage medium.
In the computer device shown in fig. 1, the network interface 1004 is mainly used for data communication with a network server; the user interface 1003 is mainly used for data interaction with a user; the processor 1001 and the memory 1005 in the computer device of the present invention may be provided in the computer device, where the computer device invokes the aircraft component assembly profile tolerance obtaining device stored in the memory 1005 through the processor 1001, and executes the aircraft component assembly profile tolerance obtaining method provided in the embodiment of the present application.
Referring to fig. 2, based on the hardware device of the foregoing embodiment, an embodiment of the present application provides an aircraft component assembly profile tolerance obtaining method, including:
s10: acquiring process assembly information of an analysis area of the appearance error of the target aircraft component; wherein, the process assembly information comprises assembly tolerance;
In the specific implementation process, carding the assembly relation of the components of the aircraft, determining an appearance error analysis area P of the components of the aircraft, and acquiring process assembly information of the area.
In this embodiment, taking a certain component 5 of an aircraft as an example, the structure is shown in fig. 10, and table 1 shows relevant parameters of process assembly information of an analysis area of the component.
TABLE 1
S20: acquiring an assembly relation chain of the analysis area according to the process assembly information; wherein the assembly relationship chain is an aircraft component assembly relationship from an analysis reference to the analysis region;
in a specific implementation process, the assembly relation chain of the analysis area is obtained according to the content in the process assembly information, the error analysis reference is associated with the existing assembly reference of the aircraft component, and is preferentially selected from the existing assembly references of the aircraft component, namely, the assembly relation of the component from the analysis reference to the analysis area.
Specifically, referring to FIG. 11, in the present embodiment, an aircraft part shape error analysis reference C is first determined 0 ，C 0 The origin is under the aircraft coordinate system (1470,11270, -702), the X-axis direction vector is (1, 0), the Y-axis direction vector is (0, 1, 0), and the Z-axis direction vector is (0, 1); as shown in fig. 6, the aircraft component assembly relationship is combed based on the process assembly information therein according to the design model or drawing, and the assembly relationship chain C from the analysis reference to the analysis area is obtained 0 -C 1 -C 2 -C 3 -C 4 -C 5 。
S30: constructing an error calculation model according to the assembly relation chain;
in the specific implementation process, a calculation model of the appearance error of the part is constructed according to the assembly relation chain. The error calculation model predicts aircraft shape errors based on existing manufacturing tolerance information.
As an optional implementation manner, the step of constructing an error calculation model according to the assembly relation chain includes:
s301: according to the assembly relation chain, a first point vector and a first normal vector of any point in the analysis area under a reference coordinate system, and a second point vector and a second normal vector of the point under the analysis area coordinate system are obtained; wherein the Z axis of the analysis region coordinate system is in the same direction as the first normal vector;
in the specific implementation process, the homogeneous coordinates of a point in the outline error analysis region P on the skeleton in the reference coordinate system C0 are set as follows Homogeneous coordinates in C5 coordinate systemIn the reference coordinate system C0 +.>Homogeneous coordinates of normal vector at point are The homogeneous coordinates are +.>
When a local coordinate system Cn is created, the Z axis and the Z axis are required to be ensuredNormal ∈10 at points>And the same direction. (the reason for the same direction should be explained)
Explanation: local coordinate system C n The outer parts (such as skins) which are positioned on the outer surfaces of the last assembled and positioned skeleton parts are tightly attached to the outer surfaces of the skeleton parts when assembled, and the manufacturing errors of the outer parts are expressed as thickness deviations and are expressed as normal variation along the outer surfaces of the skeleton. The method requires the Z axis to be ensuredNormal ∈10 at points>In the same direction, the aim is to simplify the construction complexity of the skeleton appearance error model, and simplify the attachment assembly error from matrix conversion to scalar accumulation along the normal vector direction of the skeleton appearance surface.
S302: acquiring a coordinate system conversion relation in a theoretical digital-analog state according to the first point vector, the first normal vector, the second point vector and the second normal vector;
in the specific implementation process, the coordinate system conversion relation under the theoretical digital-analog state can be obtained according to the coordinate information of the points.
As an optional implementation manner, the step of obtaining the coordinate system conversion relationship in the theoretical digital-analog state according to the first point vector, the first normal vector, the second point vector and the second normal vector includes:
the coordinate system conversion relation is obtained through the following relation:
Referring to fig. 5 and 6, the assembly is composed of a skeleton part and an external part, part 1 is the first skeleton part of the assembly, part 2 is positioned and assembled on part 1, and so on, part 3 is positioned and assembled on the last part, part 4 is the shape of the skeleton external part 3, wherein,is a first point vector whose coordinate value is +.> Is a second point vector whose coordinate value is +.> Is a first normal vector whose coordinate value is +.> Is a first normal vector whose coordinate value is +.> Represent C 1 Coordinate system to C 0 Conversion relation homogeneous matrix of coordinate system, +.>Represent C 2 Coordinate system to C 1 Conversion relation homogeneous matrix of coordinate system, +.>Represent C 3 Coordinate system to C 2 Conversion relation homogeneous matrix of coordinate system, +.>Represent C n Coordinate system to C n-1 And (5) a homogeneous matrix of the conversion relation of the coordinate system.
In the specific implementation process, according to the coordinate information of the points, the coordinate system conversion relation under the theoretical digital-analog state can be obtained:
wherein,represent C 0 Coordinate system to C 1 Conversion relation homogeneous matrix of coordinate system, +.>Represent C 1 Coordinate system to C 2 Conversion relation homogeneous matrix of coordinate system, +.>Represent C 3 Coordinate system to C 2 Conversion relation homogeneous matrix of coordinate system, +.>Represent C 4 Coordinate system to C 3 Conversion relation homogeneous matrix of coordinate system, +. >Represent C 5 Coordinate system to C 4 And (5) a homogeneous matrix of the conversion relation of the coordinate system.
Local coordinate system C 5 At the time of creation, its Z axis is aligned withNormal ∈10 at points>And the same direction.
S303: respectively constructing an aircraft component appearance standard calculation model and an aircraft component appearance error calculation model according to the coordinate system conversion relation;
in the specific implementation process, the coordinate system is converted according to the coordinate system conversion relation in the theoretical digital-analog state, and an aircraft component appearance standard calculation model and an aircraft component appearance error calculation model are constructed and are used for calculating an error value of an aircraft appearance.
As an optional implementation manner, the step of respectively constructing an aircraft component shape standard calculation model and an aircraft component shape error calculation model according to the coordinate system conversion relation includes:
constructing the aircraft component appearance standard calculation model by the following relation:
wherein ε i Represents the theoretical thickness of the outer skin 9 of the skeleton of the aircraft component,Represent C 1 Coordinate system to C 0 Conversion relation homogeneous matrix of coordinate system, +.>Represent C 2 Coordinate system to C 1 Conversion relation homogeneous matrix of coordinate system, +.>Represent C 3 Coordinate system to C 2 Coordinate systemIs a homogeneous matrix of conversion relations- >Represent C n Coordinate system to C n-1 A homogeneous matrix of the conversion relation of the coordinate system;
the aircraft component shape error calculation model is constructed by the following relation:
wherein ε' i Representing the actual thickness, delta, of the outer skin 9 (m layers) of the skeleton of the aircraft component n Representative ofError matrix of (2),Represent C 1 Coordinate system to C 0 Conversion relation homogeneous matrix of coordinate system, +.>Represent C 2 Coordinate system to C 1 Conversion relation homogeneous matrix of coordinate system, +.>Represent C 3 Coordinate system to C 2 Conversion relation homogeneous matrix of coordinate system, +.>Represent C n Coordinate system to C n-1 A homogeneous matrix of the conversion relation of the coordinate system;
in the specific implementation process, the aircraft skeleton is provided with 3 layers of outer skins 9 with theoretical thicknesses of epsilon respectively 1 ＝1.375mm，ε 2 ＝1.0mm，ε 2 ＝2.0mm。
N can be calculated according to the model i ＝[780.105 -95.032 812.129 1] T 。
Constructing an aircraft component shape calculation model including part manufacturing errors and assembly errors:
wherein ε' 1 Represents the actual thickness value of the spacer d7, ε' 2 Represents the actual thickness value of the gasket e8, ε' 3 Representing the actual thickness value of the outer skin 9, the values are:
specifically, in the aircraft component shape standard calculation model and the aircraft component shape error calculation model, Δδ n Representative ofAssembling an error matrix of the relation matrix, the error matrix being similar to:
Referring to FIG. 7, wherein u n 、v n 、w n 、α n 、β n 、γ n For the error matrix parameters, an error matrix parameter vector X is formed i ＝(u i 、v i 、w i 、α i 、β i 、γ i ) I=1, 2, n, the specific value of the vector depends on the characteristic form, the assembly positioning form and the tolerance value t i 。
Error matrix parameter vector X i And tolerance t i Relationships between, e.gThe following steps:
when the aircraft component is contoured as a planar positioning feature, reference is made to figure 8,t i Representing the planar feature tolerance, L1 and L2 represent the long-side and short-side values of the planar feature region, respectively, and the error matrix parameter vector is:
when the aircraft component is shaped as a cylindrical bit feature, t i Representing the tolerance of the plane feature, and H represents the length of the cylindrical feature respectively, wherein the error matrix parameter vector is as follows:
in the present embodiment, Δδ 1 ～Δδ 5 Respectively represent an assembly gateway matrixThe error matrix of (a) is: />
Wherein u is i 、v i 、w i 、α i 、β i 、γ i For error matrix parameters, corresponding to assembly or form and position tolerances t i ，C 1 And C 2 For plane fit, C 3 And C 4 For plane fit, C 5 The local features are approximate facets, and the specific structure of the error matrix parameter vectorThe method comprises the following steps:
/>
s304: and constructing the error calculation model according to the aircraft component appearance standard calculation model and the aircraft component appearance error calculation model.
In a specific implementation, an error calculation model at an aircraft part appearance point Ni is constructed according to an aircraft part appearance standard calculation model and an aircraft part appearance error calculation model, and the error calculation model is used for subsequently identifying errors at the point.
As an alternative embodiment, the step of constructing the error calculation model according to the aircraft component shape standard calculation model and the aircraft component shape error calculation model includes:
the error calculation model is constructed by the following relation:
wherein M is i Model N for calculating shape error of aircraft component i Is the appearance standard of the aircraft partsAnd calculating a model.
In the specific implementation process, constructing an appearance error calculation model at an appearance point Ni of the aircraft component as follows:wherein (1)>As a first normal vector, M i Model N for calculating shape error of aircraft component i A model is calculated for the aircraft part form standard.
S40: obtaining a prediction index of the appearance error of the target aircraft component according to the error calculation model;
in the specific implementation process, the prediction index refers to the average value and standard deviation of data in the error calculation data set, and is used for judging the subsequent qualification rate and the like. The prediction method is suitable for scenes similar to manufacturing error prediction or process tolerance design, and has good popularization value.
As an optional embodiment, the step of obtaining a prediction index of the shape error of the target aircraft component according to the error calculation model includes:
S401: obtaining an error matrix according to the assembly tolerance;
in the concrete implementation process, according to the assembly tolerance t i For error matrix parameter vector { X i I=1, 2,.. i I=1, 2, & gt, n }, the sampling method is that error parameters are sampled according to normal distribution, namelyIn the present embodiment, the error matrix parameter vector is { X } i I=1, 2,..5 }, the error parameter tolerance range is 6 standard deviations of the normal distribution, namely: />
/>
According toConversion to obtain an error matrix { delta } i ，i＝1，2，...，5}。
S402: acquiring the actual thickness of the outer skin of the skeleton of the aircraft component;
in a specific embodiment, the actual thickness value ε 'of the outer skin 9 is measured for aircraft component errors' i The data sampling is carried out by sampling the actual thickness parameters of the outer skin 9 according to normal distribution, namelyThe sampling standard deviation is 1/6 of the tolerance range of the error parameter, namely +.>
Specifically, in this embodiment, data sampling is performed according to the above method, and the result is:
ε′ 1 ～N(1.375，0.11/6)
ε′ 2 ～N(1，0.08/6)
ε′ 3 ～N(2，0.16/6)
s403: based on the error matrix and the actual thickness of the skin, performing simulation calculation on an aircraft component shape error calculation model to obtain an aircraft component shape error simulation data set;
In a specific implementation process, a Monte Carlo method is adopted, an error matrix is obtained according to steps S401 and S402, and simulation calculation is performed on the aircraft component shape error calculation model obtained in step S303, so as to obtain an aircraft component shape simulation data set { M } i I=1, 2,..and N }, wherein, N is the sampling calculation times of Monte Carlo simulation. The Monte Carlo method is adopted for simulation calculation, and the calculation result has statistical significance and good robustness. In this embodiment, 10000 times are taken as sampling calculation times of the monte carlo simulation.
S404: obtaining an error calculation data set according to the aircraft component shape error simulation data set and the error calculation model;
in a specific implementation, the data set M is simulated according to the shape of the aircraft component i And error calculation model E i An error calculation dataset is obtained. In the present embodiment, the error calculation value set at the Ni point on the aircraft component is { E } i |E i ∈E，i＝1，2，...，10000}。
S405: calculating a data set according to the error, and acquiring the prediction index; and calculating the average value and standard deviation of the data in the data set for the error by the prediction index.
In the specific implementation process, the prediction index refers to the average value and standard deviation of data in the error calculation data set. From dataset { E i Mean μ, standard deviation σ, maximum E max Minimum value E min Median E min . Thus, a prediction index of the shape error at the Ni point on the aircraft component is obtained.
In this embodiment, the prediction index related data of the appearance error at the upper Ni point of a certain part 5 of the aircraft to be measured is: form error mean μ= -0.073, form error standard deviation σ = 0.226, form error maximum E max =0.923, shape error minimum E min = -0.949, median of shape error E mid ＝-0.070。
S50: and acquiring the assembly appearance tolerance value of the target aircraft component according to the prediction index.
In the specific implementation process, the tolerance qualification rate of the manufacturing process is calculated according to the prediction index of the appearance error of the part, and the assembly appearance tolerance value of the target aircraft part is obtained according to the tolerance qualification rate of the manufacturing process meeting the manufacturing requirement.
As an optional embodiment, the step of obtaining the fitting shape tolerance value of the target aircraft component according to the prediction index includes:
judging whether the manufacturing tolerance qualification rate of the assembly tolerance is more than or equal to the preset process requirement; wherein, the manufacturing tolerance qualification rate is obtained according to the prediction index;
If yes, the assembly tolerance is the assembly external tolerance value;
if not, the assembly tolerance corresponding to the error parameter with the largest contribution degree in the error matrix is selected for adjustment, so that the assembly appearance tolerance value is obtained.
In the specific implementation process, the tolerance qualification rate C of the manufacturing process is calculated according to the statistical index of the calculated value of the component shape error pk Judging whether the manufacturing tolerance qualification rate is greater than or equal to a preset process requirement U, wherein the U value is determined by design or process requirements, and under the general condition, U is more than or equal to 1: when C pk When the tolerance parameter is not less than U, judging that the tolerance parameter in the current manufacturing process can meet the requirement of the appearance tolerance of the aircraft component, wherein the assembly tolerance at the moment is the assembly appearance tolerance value; when C pk When the number is less than U, judging that the manufacturing is performed currentlyCheng Rongcha parameters cannot meet the requirements of the appearance tolerance of the aircraft component, and the tolerance parameters need to be adjusted, namely, the assembly tolerance corresponding to the error parameter with the largest contribution degree in the error matrix is selected for adjustment, so as to obtain the assembly appearance tolerance value.
Specifically, in the present embodiment, C pk =0.629, U is 1, c pk < U, it is known that the current manufacturing process tolerance parameters need to be adjusted.
As an alternative embodiment, the manufacturing tolerance yield is obtained by the following relation:
Wherein T is max Is the upper limit value of the appearance error of the aircraft component, T min The lower limit value, mu and sigma of the appearance error of the aircraft component are used as prediction indexes, mu is an average value,sigma is standard deviation>
In a specific implementation process, the manufacturing tolerance qualification rate in this embodiment is:
as an alternative embodiment, the contribution is obtained by the following relation:
wherein t is i For assembly tolerances, σ is the standard deviation,
in the specific implementation process, the contribution degree of each error parameter is calculated according to the statistical indexThe calculation method comprises the following steps: />The calculation results of the contribution degree are shown in table 2:
TABLE 2
At this time, if not, a step of selecting an assembly tolerance corresponding to an error parameter with the largest contribution degree in the error matrix to adjust to obtain the assembly appearance tolerance value, including:
and modifying the tolerance zone position and the tolerance zone width of the assembly tolerance corresponding to the error parameter with the largest contribution degree to obtain the assembly external tolerance value.
In the specific implementation process, if the manufacturing tolerance qualification rate of the assembly tolerance is smaller than the preset process requirement, the assembly tolerance corresponding to the error parameter with the largest contribution degree in the error matrix is selected for adjustment, and the steps S40 and S50 are repeated until the obtained C pk The values meet the U-value requirement. The specific adjustment method is as follows:
1. selecting an assembly tolerance t corresponding to the error parameter with the largest contribution degree i ；
2. Adjusting and modifying the position of the tolerance zone of the assembly tolerance, and re-acquiring the error parameter X i And epsilon' 1 Obtaining the prediction index: average value ofStandard deviation->Repeating the steps S40 and S50 to calculate the appearance error symmetry index E sym Until the following conditions are satisfied:
the index is used to adjust the tolerance zone position of the assembly tolerance.
3. Adjusting the tolerance band width of the modified assembly tolerance so that the recalculated C pk The values meet the U-value requirement.
During the adjustment process, the assembly tolerance can be increased or decreased by an increment value, the adjustment increment is 0.001mm, but the requirement is that: assembly tolerance t i Standard deviation of corresponding generated error parameters
Generally, any manufacturing tolerance has an upper limit and a lower limit, and the smaller the upper limit and the lower limit, the narrower the representative tolerance range, which means that the higher the manufacturing accuracy requirement is, the more manufacturing cost is correspondingly paid. The above limitation is to control a certain assembly tolerance range not to be too small, so as to limit the control of the manufacturing precision not to be too severe, and save the manufacturing cost as much as possible.
In this embodiment, according to the result of contribution calculation, the assembly tolerance t 1 、t 3 、t 5 Adjusting, repeating the steps S40 and S50, recalculating the appearance error statistical index, and recalculating the tolerance qualification rate C of the manufacturing process pk Whether the requirement of not less than U is met or not, and finally, after the adjustment and optimization by the steps, error parameters are shown in the table 3:
TABLE 3 Table 3
The outline error prediction index related data at this time is: the mean value μ=0.001, the standard deviation σ=0.166, and the maximum value E max =0.634, the shape error is the mostSmall value E min = -0.649, median of shape error E mid =0.002. Appearance error symmetry index E at this time sym =0.995, manufacturing process tolerance yield C pk =1.005, satisfy C pk And the requirement of U is not less than.
Adjusted assembly tolerance t i Is the final assembly appearance tolerance value { S } meeting the appearance error requirement of the aircraft component i }。
It should be understood that the foregoing is merely illustrative, and the technical solutions of the present application are not limited in any way, and those skilled in the art may perform the setting based on the needs in practical applications, and the present application is not limited herein.
Through the description, it is easy to find that the method is based on the real assembly relation of the airplane, and takes the manufacturing error and the assembly error of the parts into consideration, and adopts a three-dimensional tolerance modeling method to predict, so that the situation of judging by experience is avoided, and the accuracy of adjustment is improved; moreover, the appearance of the aircraft component is predicted by adopting a simulation calculation method, so that a tool for measuring and judging the real appearance of the aircraft is eliminated, the cost is low, and the efficiency is high; furthermore, the aircraft appearance error can be predicted based on the existing manufacturing tolerance design error model, and the manufacturing process tolerance can be designed based on the target aircraft appearance error requirement, so that the processes of repeated trial and error and adjustment of the traditional method are avoided, and the quality and efficiency are both considered.
Referring to fig. 3, based on the same inventive concept, an embodiment of the present application further provides an apparatus for obtaining an assembly profile tolerance of an aircraft component, including:
the process assembly information acquisition module is used for acquiring process assembly information of an analysis area of the appearance error of the target aircraft component; wherein, the process assembly information comprises assembly tolerance;
the assembly relation chain acquisition module is used for acquiring the assembly relation chain of the analysis area according to the process assembly information; wherein the assembly relationship chain is an aircraft component assembly relationship from an analysis reference to the analysis region;
the error calculation model construction module is used for constructing an error calculation model according to the assembly relation chain;
the prediction index acquisition module is used for acquiring a prediction index of the appearance error of the target aircraft component according to the error calculation model;
and the tolerance value acquisition module is used for acquiring the assembly appearance tolerance value of the target aircraft component according to the prediction index.
It should be noted that, each module in the aircraft component assembly profile tolerance obtaining apparatus in this embodiment corresponds to each step in the aircraft component assembly profile tolerance obtaining method in the foregoing embodiment, so the specific implementation manner of this embodiment may refer to the implementation manner of the foregoing aircraft component assembly profile tolerance obtaining method, and will not be repeated herein.
Furthermore, in an embodiment, the embodiments of the present application also provide a computer device, the device comprising a processor, a memory and a computer program stored in the memory, which when executed by the processor, implements the steps of the method in the previous embodiments.
Furthermore, in an embodiment, the embodiments of the present application further provide a computer storage medium, on which a computer program is stored, which when being executed by a processor, implements the steps of the method in the previous embodiments.
In some embodiments, the computer readable storage medium may be FRAM, ROM, PROM, EPROM, EEPROM, flash memory, magnetic surface memory, optical disk, or CD-ROM; but may be a variety of devices including one or any combination of the above memories. The computer may be a variety of computing devices including smart terminals and servers.
In some embodiments, the executable instructions may be in the form of programs, software modules, scripts, or code, written in any form of programming language (including compiled or interpreted languages, or declarative or procedural languages), and they may be deployed in any form, including as stand-alone programs or as modules, components, subroutines, or other units suitable for use in a computing environment.
As an example, the executable instructions may, but need not, correspond to files in a file system, may be stored as part of a file that holds other programs or data, for example, in one or more scripts in a hypertext markup language (HTML, hyper Text Markup Language) document, in a single file dedicated to the program in question, or in multiple coordinated files (e.g., files that store one or more modules, sub-programs, or portions of code).
As an example, executable instructions may be deployed to be executed on one computing device or on multiple computing devices located at one site or, alternatively, distributed across multiple sites and interconnected by a communication network.
It should be noted that, in this document, the terms "comprises," "comprising," or any other variation thereof, are intended to cover a non-exclusive inclusion, such that a process, method, article, or system that comprises a list of elements does not include only those elements but may include other elements not expressly listed or inherent to such process, method, article, or system. Without further limitation, an element defined by the phrase "comprising one … …" does not exclude the presence of other like elements in a process, method, article, or system that comprises the element.
The foregoing embodiment numbers of the present application are merely for describing, and do not represent advantages or disadvantages of the embodiments.
From the above description of the embodiments, it will be clear to those skilled in the art that the above-described embodiment method may be implemented by means of software plus a necessary general hardware platform, but of course may also be implemented by means of hardware, but in many cases the former is a preferred embodiment. Based on such understanding, the technical solution of the present application may be embodied essentially or in a part contributing to the prior art in the form of a software product stored in a storage medium (e.g. read-only memory/random-access memory, magnetic disk, optical disk), comprising several instructions for causing a multimedia terminal device (which may be a mobile phone, a computer, a television receiver, or a network device, etc.) to perform the method described in the embodiments of the present application.
The foregoing description is only of the preferred embodiments of the present application, and is not intended to limit the scope of the claims, and all equivalent structures or equivalent processes using the descriptions and drawings of the present application, or direct or indirect application in other related technical fields are included in the scope of the claims of the present application.

Claims (11)

1. A method of obtaining an assembly form tolerance of an aircraft component, comprising the steps of:
acquiring process assembly information of an analysis area of the appearance error of the target aircraft component; wherein, the process assembly information comprises assembly tolerance;
acquiring an assembly relation chain of the analysis area according to the process assembly information; wherein the assembly relationship chain is an aircraft component assembly relationship from an analysis reference to the analysis region;
constructing an error calculation model according to the assembly relation chain;
obtaining a prediction index of the appearance error of the target aircraft component according to the error calculation model;
and acquiring the assembly appearance tolerance value of the target aircraft component according to the prediction index.
2. The aircraft component assembly profile tolerance acquisition method of claim 1, wherein the step of constructing an error calculation model from the assembly relationship chain comprises:
according to the assembly relation chain, a first point vector and a first normal vector of any point in the analysis area under a reference coordinate system, and a second point vector and a second normal vector of the point under the analysis area coordinate system are obtained; wherein the Z axis of the analysis region coordinate system is in the same direction as the first normal vector;
Acquiring a coordinate system conversion relation in a theoretical digital-analog state according to the first point vector, the first normal vector, the second point vector and the second normal vector;
respectively constructing an aircraft component appearance standard calculation model and an aircraft component appearance error calculation model according to the coordinate system conversion relation;
and constructing the error calculation model according to the aircraft component appearance standard calculation model and the aircraft component appearance error calculation model.
3. The aircraft component assembly profile tolerance acquisition method according to claim 2, wherein the step of acquiring the coordinate system conversion relationship in the theoretical digital-analog state based on the first point vector, the first normal vector, the second point vector, and the second normal vector comprises:
the coordinate system conversion relation is obtained through the following relation:
wherein,is a first point vector whose coordinate value is +.> Is a second point vector whose coordinate value is +.> Is a first normal vector with coordinate values of Is a first normal vector whose coordinate value is +.> Represent C 1 Coordinate system to C 0 Conversion relation homogeneous matrix of coordinate system, +.>Represent C 2 Coordinate system to C 1 Conversion relation homogeneous matrix of coordinate system, +. >Represent C 3 Coordinate system to C 2 Conversion relation homogeneous matrix of coordinate system, +.>Represent C n Coordinate system to C n-1 And (5) a homogeneous matrix of the conversion relation of the coordinate system.
4. The aircraft component assembly profile tolerance acquisition method according to claim 2, wherein the step of constructing an aircraft component profile standard calculation model and an aircraft component profile error calculation model, respectively, based on the coordinate system conversion relationship, comprises:
constructing the aircraft component appearance standard calculation model by the following relation:
wherein ε i Represents the theoretical thickness of the outer skin of the skeleton of the aircraft component,Represent C 1 Coordinate system to C 0 Conversion relation homogeneous matrix of coordinate system, +.>Represent C 2 Coordinate system to C 1 Conversion relation homogeneous matrix of coordinate system, +.>Represent C 3 Coordinate system to C 2 Conversion relation homogeneous matrix of coordinate system, +.>Represent C n Coordinate system to C n-1 A homogeneous matrix of the conversion relation of the coordinate system;
the aircraft component shape error calculation model is constructed by the following relation:
wherein ε i ′ Representing the actual thickness of the skin (m layers) outside the skeleton of an aircraft component, delta n Representative ofError matrix of +.>Represent C 1 Coordinate system to C 0 The transformation relation of the coordinate system is homogeneous matrix,/>represent C 2 Coordinate system to C 1 Conversion relation homogeneous matrix of coordinate system, +. >Represent C 3 Coordinate system to C 2 Conversion relation homogeneous matrix of coordinate system, +.>Represent C n Coordinate system to C n-1 A homogeneous matrix of the conversion relation of the coordinate system;
the step of constructing the error calculation model according to the aircraft component appearance standard calculation model and the aircraft component appearance error calculation model comprises the following steps:
the error calculation model is constructed by the following relation:
wherein M is i Model N for calculating shape error of aircraft component i A model is calculated for the aircraft part form standard.
5. The aircraft component assembly profile tolerance acquisition method according to claim 1, wherein the step of acquiring a predictor of the target aircraft component profile error from the error calculation model comprises:
obtaining an error matrix according to the assembly tolerance;
acquiring the actual thickness of the outer skin of the skeleton of the aircraft component;
based on the error matrix and the actual thickness of the skin, performing simulation calculation on an aircraft component shape error calculation model to obtain an aircraft component shape error simulation data set;
obtaining an error calculation data set according to the aircraft component shape error simulation data set and the error calculation model;
calculating a data set according to the error, and acquiring the prediction index; and calculating the average value and standard deviation of the data in the data set for the error by the prediction index.
6. The aircraft component assembly profile tolerance acquisition method according to claim 1, wherein the step of acquiring the assembly profile tolerance value of the target aircraft component based on the prediction index comprises:
judging whether the manufacturing tolerance qualification rate of the assembly tolerance is more than or equal to the preset process requirement; wherein, the manufacturing tolerance qualification rate is obtained according to the prediction index;
if yes, the assembly tolerance is the assembly external tolerance value;
if not, the assembly tolerance corresponding to the error parameter with the largest contribution degree in the error matrix is selected for adjustment, so that the assembly appearance tolerance value is obtained.
7. The aircraft component assembly profile tolerance acquisition method of claim 6, wherein the manufacturing tolerance yield is obtained by the relationship:
wherein T is max Is the upper limit value of the appearance error of the aircraft component, T min The lower limit value, mu and sigma of the appearance error of the aircraft component are used as prediction indexes, mu is an average value,sigma is standard deviation>
8. The aircraft component assembly profile tolerance acquisition method according to claim 6, wherein the contribution is obtained by the following relation:
wherein t is i For assembly tolerances, σ is the standard deviation,
If not, selecting an assembly tolerance corresponding to an error parameter with the largest contribution degree in the error matrix for adjustment to obtain the assembly external tolerance value, wherein the step comprises the following steps:
and modifying the tolerance zone position and the tolerance zone width of the assembly tolerance corresponding to the error parameter with the largest contribution degree to obtain the assembly external tolerance value.
9. An aircraft component assembly profile tolerance acquisition device, comprising:
the process assembly information acquisition module is used for acquiring process assembly information of an analysis area of the appearance error of the target aircraft component; wherein, the process assembly information comprises assembly tolerance;
the assembly relation chain acquisition module is used for acquiring the assembly relation chain of the analysis area according to the process assembly information; wherein the assembly relationship chain is an aircraft component assembly relationship from an analysis reference to the analysis region;
the error calculation model construction module is used for constructing an error calculation model according to the assembly relation chain;
the prediction index acquisition module is used for acquiring a prediction index of the appearance error of the target aircraft component according to the error calculation model;
and the tolerance value acquisition module is used for acquiring the assembly appearance tolerance value of the target aircraft component according to the prediction index.
10. A computer device, characterized in that it comprises a memory in which a computer program is stored and a processor which executes the computer program, implementing the method according to any of claims 1-8.
11. A computer readable storage medium, having stored thereon a computer program, the computer program being executable by a processor to implement the method of any of claims 1-8.
CN202311164503.8A 2023-09-11 2023-09-11 Method, device, equipment and storage medium for acquiring assembly shape tolerance of aircraft component Pending CN117349955A (en)

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