CN112035959B - Method for rapidly determining airfoil rigid shaft by using finite element analysis result of whole machine - Google Patents

Method for rapidly determining airfoil rigid shaft by using finite element analysis result of whole machine Download PDF

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CN112035959B
CN112035959B CN202010910989.5A CN202010910989A CN112035959B CN 112035959 B CN112035959 B CN 112035959B CN 202010910989 A CN202010910989 A CN 202010910989A CN 112035959 B CN112035959 B CN 112035959B
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李伟
闫雨哲
邢华璐
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Shenyang Aircraft Design and Research Institute Aviation Industry of China AVIC
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Abstract

The application belongs to the field of airplane strength design, and particularly relates to a method for quickly determining an airfoil rigid shaft by using a complete machine finite element analysis result. The method comprises the following steps: the method comprises the following steps: obtaining a complete machine finite element model, and selecting 3 loading points on an airfoil end rib of the complete machine finite element model; step two: obtaining coordinates of the 3 loading points, and determining loads for loading each loading point, wherein the loads for loading the 3 loading points are equal; step three: applying constraint on the wing root rib of the full-machine finite element model, respectively applying loads to the 3 loading points, carrying out finite element numerical calculation once during each loading, and obtaining a torsion angle of the wing end rib relative to the wing root rib when the corresponding loading point is loaded; step four: and establishing a functional relation between the torsion angle and the coordinates of the loading points, substituting the torsion angle and the coordinates of the 3 loading points to obtain three equation sets, and solving to obtain the position of the airfoil rigid shaft.

Description

Method for rapidly determining airfoil rigid shaft by using finite element analysis result of whole machine
Technical Field
The application belongs to the field of airplane strength design, and particularly relates to a method for quickly determining an airfoil rigid shaft by using a complete machine finite element analysis result.
Background
The wings are the main components of an aircraft, the main function of which is to generate lift under the action of airflow. Under the combined action of aerodynamic load and inertial load, the wing generates bending deformation and torsion. The bending rigidity and the torsional rigidity of the airfoil structure influence the deformation of the airplane in the flying process, and the higher the rigidity is, the smaller the deformation is, the lower the rigidity is, the larger the deformation is, and the more obvious the influence on the aerodynamic performance is. The bending stiffness is generally determined by the upper and lower skins of the airfoil, the torsional stiffness is determined by the closed chamber formed by the upper and lower skins and the skeleton structure, and the torque is determined by the position of a load action point relative to a rigid shaft. The rigid shaft is a rigid-center connecting line of each tangent plane of the wing structure, the load acts on the rigid shaft, the wing does not generate a torsion angle, the position of the typical wing rigid shaft is determined for the wing, and the position is shown in figure 1. After the rigid shaft position is determined in the scheme stage, the position of the rigid shaft is ensured to be not changed greatly through parameter design in the detailed design process, so that the position of the wing rigid shaft needs to be evaluated quickly and accurately.
Accordingly, a technical solution is desired to overcome or at least alleviate at least one of the above-mentioned drawbacks of the prior art.
Disclosure of Invention
The application aims to provide a method for rapidly determining an airfoil rigid shaft by using a full-machine finite element analysis result so as to solve at least one problem in the prior art.
The technical scheme of the application is as follows:
a method for rapidly determining an airfoil rigid shaft by using a full-machine finite element analysis result comprises the following steps:
the method comprises the following steps: acquiring a full-machine finite element model, and selecting 3 loading points on a wing end rib of the full-machine finite element model;
step two: obtaining coordinates of the 3 loading points, and determining loads for loading each loading point, wherein the loads for loading the 3 loading points are equal;
step three: applying constraint on the wing root rib of the full-machine finite element model, respectively applying load to the 3 loading points, and performing finite element numerical calculation once for each loading to obtain a torsion angle of the wing end rib relative to the wing root rib when the corresponding loading point is loaded;
step four: establishing a functional relationship between the torsion angle and the loading point coordinate, including:
for the loading point (x) i ，y i ) When a load P is applied, the torque T generated By the load P relative to the airfoil rigid shaft Ax + By + C is 0 i The expression is:
Figure BDA0002663260410000021
at the torque T i Under the action of the torsion angle of the wing end rib relative to the wing root
Figure BDA0002663260410000024
Expression ofThe formula is as follows:
Figure BDA0002663260410000022
wherein, the two expressions are combined to obtain:
Figure BDA0002663260410000023
wherein l i Is a loading point (x) i ，y i ) Distance from the wing rigid shaft, G is inherent torsional rigidity of the wing, and A, B, C is a parameter of a wing rigid shaft equation;
and substituting the torsion angle and the coordinates of the 3 loading points to obtain three equation sets, solving, and calculating three parameters A, B, C in the expression of the airfoil rigid shaft to obtain the position of the airfoil rigid shaft.
Optionally, in step two, when coordinates of the 3 loading points are acquired, a coordinate system is defined by taking a wing root rib leading edge of the full-machine finite element model as a coordinate origin.
Optionally, in step three, the wing root rib of the full-machine finite element model is constrained by a fixing bracket.
Optionally, in the third step, when a corresponding loading point is obtained through finite element numerical calculation and loading is performed, in the torsion angle of the wing end rib relative to the wing root, the torsion angle of the wing end rib relative to the root under the action of the load P is obtained through Y-direction displacement calculation according to the front point and the rear point of the wing end rib.
The invention has at least the following beneficial technical effects:
according to the method for rapidly determining the wing rigid shaft by using the finite element analysis result of the whole aircraft, the node force is applied to the finite element model of the whole aircraft, the relation between the node force and the torsion angle can be conveniently obtained, the complex tangent plane rigidity characteristic calculation is avoided, the tangent plane rigidity is the inherent attribute of the tangent plane, the complex solution is not carried out, the wing rigid shaft position is simply and conveniently obtained by solving the equation set, the calculation efficiency is improved, a basis is provided for the control of the rigid shaft and the torsional rigidity design in the detailed design process of the aircraft, and a basis is provided for optimizing the wing scheme.
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FIG. 1 is a schematic representation of a typical prior art wing rigid shaft location;
FIG. 2 is a flow chart of a method for rapidly determining an airfoil rigid shaft using full-machine finite element analysis results according to an embodiment of the present application.
Detailed Description
In order to make the implementation objects, technical solutions and advantages of the present application clearer, the technical solutions in the embodiments of the present application will be described in more detail below with reference to the drawings in the embodiments of the present application. In the drawings, the same or similar reference numerals denote the same or similar elements or elements having the same or similar functions throughout. The described embodiments are a subset of the embodiments in the present application and not all embodiments in the present application. The embodiments described below with reference to the drawings are exemplary and intended to be used for explaining the present application and should not be construed as limiting the present application. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present application. Embodiments of the present application will be described in detail below with reference to the accompanying drawings.
In the description of the present application, it is to be understood that the terms "center", "longitudinal", "lateral", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", and the like indicate orientations or positional relationships based on those shown in the drawings, and are used merely for convenience in describing the present application and for simplifying the description, and do not indicate or imply that the referenced device or element must have a particular orientation, be constructed in a particular orientation, and be operated, and therefore should not be construed as limiting the scope of the present application.
The present application is described in further detail below with reference to fig. 1-2.
The application provides a method for rapidly determining an airfoil rigid shaft by using a complete machine finite element analysis result, which comprises the following steps:
the method comprises the following steps: obtaining a complete machine finite element model, selecting 3 loading points on an airfoil end rib of the complete machine finite element model, and selecting 3 loading points at any position of the airfoil end rib;
step two: acquiring coordinates of 3 loading points, and determining loads for loading each loading point, wherein the loads for loading the 3 loading points are equal;
step three: applying constraint on the wing root rib of the full-aircraft finite element model, respectively applying load to 3 loading points, performing finite element numerical calculation once for each loading, and obtaining torsion angles of the end rib of the aircraft wing relative to the wing root rib when the corresponding loading point is loaded, namely obtaining 3 torsion angles corresponding to the loading points;
step four: establishing a functional relation between a torsion angle and a loading point coordinate, wherein the relation between the torsion angle and the loading point applying a load is completely determined by the applied load and the inherent torsional rigidity of the wing, and the process of establishing the functional relation between the torsion angle and the loading point coordinate comprises the following steps of by utilizing the inherent physical characteristics of the torsional rigidity of the wing surface structure and the rigid shaft of the wing surface:
to the loading point (x) i ，y i ) When a load P is applied, the torque T generated By the load P relative to the airfoil rigid shaft Ax + By + C is 0 i The expression is:
Figure BDA0002663260410000041
at a torque T i Under the action of the torsion angle of the wing end rib relative to the wing root
Figure BDA0002663260410000042
The expression is as follows:
Figure BDA0002663260410000043
wherein, the two expressions are combined to obtain:
Figure BDA0002663260410000044
wherein l i Is a loading point (x) i ，y i ) Distance from the wing rigid shaft, G is inherent torsional rigidity of the wing, and A, B, C is a parameter of a wing rigid shaft equation; the inherent torsional rigidity G of the wing is a fixed value under the dimensional parameters determined by the wing, and the value of the inherent torsional rigidity G is independent of the loading point.
And substituting the torsion angle and the coordinates of the 3 loading points to obtain three equation sets, solving, and calculating three parameters A, B, C in the expression of the airfoil rigid shaft to obtain the position of the airfoil rigid shaft.
In one embodiment of the present application, in step two, when acquiring the coordinates of the 3 loading points, it is preferable to define a coordinate system with the leading edge of the airfoil root rib of the full finite element model as the origin of coordinates.
In one embodiment of the present application, in step three, the wing root rib of the full-aircraft finite element model is constrained by a fixed bracket. When the corresponding loading point is obtained through finite element numerical calculation and loading is carried out, the torsion angle of the wing end rib relative to the wing root comprises the following steps: firstly, applying constraint on a wing root rib in a full-aircraft finite element model; loading a load P at a selected load application point; submitting operation; and then reading a displacement calculation result, and calculating to obtain a torsion angle of the wing end rib relative to the wing root under the action of the load P according to the Y-direction displacement of the front point and the rear point of the wing end rib.
According to the method for rapidly determining the airfoil rigid shaft by using the finite element analysis result of the whole machine, in the previous step, coordinates of 3 loading points are obtained, wherein the coordinates are (x) 1 ，y 1 )、(x 2 ，y 2 )、(x 3 ，y 3 ) Determining the load P for loading each loading point, and determining the torsion angle of the wing end rib relative to the wing root rib when the corresponding 3 loading points are loaded
Figure BDA0002663260410000051
Then, establishing a functional relation between the torsion angle and the coordinates of the loading points, and substituting the torsion angle and the coordinates of the 3 loading points to obtain three equation sets:
Figure BDA0002663260410000052
Figure BDA0002663260410000053
Figure BDA0002663260410000054
and finally, solving, and calculating three parameters A, B, C in the expression of the airfoil rigid shaft to obtain the position of the airfoil rigid shaft.
In one embodiment of the application, in the scheme of establishing an equation set by using a numerical test and solving to determine the rigid shaft of the airfoil, a coordinate system with an origin as a leading edge of a wing root rib is defined, coordinates of wing root rib points under the coordinate system are (0, 0), (0, -3000), and coordinates of wing end rib points are (5000, -2505), (5000, -4000); defining an airfoil rigid shaft formula as Ax + By + C being 0 under a coordinate system; in this embodiment, the node positions are changed by three load points (5000, -2505), (5000, -2880), (5000, -3255), the load P loaded on each load point is 10000N, and 3 corresponding torsion angles are obtained, namely 0.0376rad, 0.092rad and 0.169 rad; establishing an equation set of a torsion angle and a loading point, and solving the equation set to obtain A, B, C values in a rigid shaft formula, wherein the values are 351, 1000 and 1500000 respectively; the value of A, B, C can be used to obtain the formula 351x +1000y +1500000 as 0. By the method, the wing surface rigid shaft position of the wing can be simply, conveniently and accurately obtained.
The method for rapidly determining the airfoil rigid shaft by using the full-machine finite element analysis result considers the design start of an airplane scheme, utilizes the full-machine finite element model to carry out force transmission analysis, can conveniently obtain the relation between node force and a torsion angle by applying node force to the full-machine finite element model, avoids complex section rigidity characteristic calculation, does not carry out complex solution because the section rigidity is inherent attribute of the section, and therefore can simply analyze the existing full-machine finite element model to obtain data and can simply, conveniently and rapidly determine the position of the airfoil rigid shaft by function calculation.
The method for rapidly determining the airfoil rigid shaft by using the finite element analysis result of the whole aircraft is simple and novel, has sufficient theoretical basis, improves the calculation efficiency, provides basis for the control of the rigid shaft and the torsional rigidity design in the detailed design process of the aircraft, and provides basis for optimizing the wing scheme.
The above description is only for the specific embodiments of the present application, but the scope of the present application is not limited thereto, and any changes or substitutions that can be easily conceived by those skilled in the art within the technical scope of the present application should be covered within the scope of the present application. Therefore, the protection scope of the present application shall be subject to the protection scope of the claims.

Claims (4)

1. A method for rapidly determining an airfoil rigid shaft by using a full-machine finite element analysis result is characterized by comprising the following steps of:
the method comprises the following steps: acquiring a full-machine finite element model, and selecting 3 loading points on a wing end rib of the full-machine finite element model;
step two: obtaining coordinates of the 3 loading points, and determining loads for loading each loading point, wherein the loads for loading the 3 loading points are equal;
step three: applying constraint on the wing root rib of the full-machine finite element model, respectively applying loads to the 3 loading points, carrying out finite element numerical calculation once during each loading, and obtaining a torsion angle of the wing end rib relative to the wing root rib when the corresponding loading point is loaded;
step four: establishing a functional relationship between the torsion angle and the coordinates of the loading point, including:
for the loading point (x) i ，y i ) When a load P is applied, the torque T generated By the load P relative to the airfoil rigid shaft Ax + By + C is 0 i The expression is:
Figure FDA0002663260400000011
at the torque T i Under the action of the torsion angle of the wing end rib relative to the wing root
Figure FDA0002663260400000012
The expression is as follows:
Figure FDA0002663260400000013
wherein, the two expressions are combined to obtain:
Figure FDA0002663260400000014
wherein l i Is a loading point (x) i ，y i ) The distance from the wing surface rigid shaft, G is the inherent torsional rigidity of the wing, and A, B, C is the parameter of the wing surface rigid shaft equation;
and substituting the torsion angle and the coordinates of the 3 loading points to obtain three equation sets, solving, and calculating three parameters A, B, C in the expression of the airfoil rigid shaft to obtain the position of the airfoil rigid shaft.
2. The method for rapidly determining the airfoil rigid shaft by using the results of the all-machine finite element analysis according to claim 1, wherein in the second step, when coordinates of 3 loading points are obtained, a coordinate system is defined by taking the leading edge of the airfoil root rib of the all-machine finite element model as a coordinate origin.
3. The method for rapidly determining the airfoil rigid shaft by using the results of the all-machine finite element analysis according to claim 1, wherein in step three, the wing root rib of the all-machine finite element model is constrained by a fixed bracket.
4. The method for quickly determining the airfoil rigid shaft by using the full-machine finite element analysis result according to claim 1, wherein in the third step, when the corresponding loading points are obtained through finite element numerical calculation and loaded, in torsion angles of the wing end rib relative to the wing root, the torsion angles of the wing end rib relative to the root under the action of the load P are obtained through Y-direction displacement calculation according to the front point and the rear point of the wing end rib.
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