CN114611208B - Aircraft composite material fuselage structure and design method thereof - Google Patents
Aircraft composite material fuselage structure and design method thereof Download PDFInfo
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Abstract
The invention discloses an aircraft composite material fuselage structure which comprises a skin, stringers and a bulkhead, wherein the skin is paved on the bulkhead, the stringers and the bulkhead are assembled to support the skin, the bulkhead comprises an outer edge strip riveted with the skin, a web plate bent and extended from the outer edge strip and an inner edge strip bent and extended from the web plate and arranged in parallel with the outer edge strip, the thickness of the inner edge strip, the web plate and the outer edge strip is the same, and the length of the inner side of the web plate is longer than the length of the outer edge strip and the inner edge strip. The invention also discloses a design method of the aircraft composite material fuselage structure. The invention has the beneficial effects that: simple structure, easy manufacture, light weight, good integrity and the like, and shows excellent mechanical properties when bearing impact load; the failure mode and load response of the airframe structure under the impact load can be accurately captured; the maneuverability is strong, and important support is provided for the design of the composite material fuselage structure.
Description
[ Field of technology ]
The invention relates to the technical field of fuselage structure design, in particular to an aircraft composite fuselage structure and a design method thereof.
[ Background Art ]
The advanced composite material structure has been widely used in the fields of automobiles, aerospace and the like because of its excellent properties such as light weight, high specific strength, high specific stiffness and the like. With the progressive maturity of composite processing technology, composite structures have progressively replaced metal structures in the aerospace field, for example, the composite usage of boeing 787 and air passenger a350 has exceeded 50%. Along with the gradual shortage of fossil energy sources such as fuel oil, light weight design is attracting more attention, and the adoption of an advanced composite material structure for the structural parts of a fuselage is of great significance for controlling fuel consumption and improving the transport capacity of an aircraft.
The overall performance of the aircraft is not only related to the high performance composite materials employed, but the design of the fuselage structural components and the connection between the structural components will have a significant impact on the performance of the aircraft. The fuselage bulkhead is used as a main structural component in the fuselage structure and is used for supporting the panel structure formed by the skin and the stringers, so that the aircraft is ensured to have enough capability of bearing external loads. The combined action of the skin, the fuselage former and the stringers is crucial for ensuring the safety of the flying craft.
A non-negligible fact is that, although aircraft manufacturing technology has matured, aircraft crashes still occur at times, which poses a great threat to occupant safety. The design of the fuselage structure enables the aircraft to show excellent crashworthiness in the event of a crash while meeting the light weight design, so that the aim of fully guaranteeing the safety of passengers is achieved, and the problem facing the aircraft manufacturing industry at present is solved. In general, the design of fuselage structures with composite materials should take into account safety, weight saving, crashworthiness, etc. In order to meet the high standards of the aircraft manufacturing industry, there is a need to propose an aircraft composite fuselage structure with excellent structural properties.
[ Invention ]
The invention discloses an aircraft composite material fuselage structure and a design method thereof, which show that the fuselage structure provided by the invention can obviously improve the mechanical property of bearing impact load by carrying out a large number of simulation experiments, thus providing necessary guarantee for passenger safety, and meanwhile, the provided damage model can effectively capture the failure mode and energy absorption characteristic of the structure and provide important reference for the design of the fuselage structure, thus effectively solving the technical problems related in the background technology.
In order to achieve the above purpose, the technical scheme of the invention is as follows:
The utility model provides an aircraft combined material fuselage structure, includes skin, stringer and bulkhead, the skin lay in on the bulkhead, the stringer with the bulkhead is furnished with and is supported the skin, the bulkhead include with skin riveted outer fringe strip, from the web that outer fringe strip buckles the extension and from the web buckle extension and with outer fringe strip parallel arrangement's inner fringe strip, inner fringe strip the web and the thickness of outer fringe strip is the same, just the inboard length of web is longer than outer fringe strip with the length of inner fringe strip.
As a preferred development of the invention, the former is of integrally formed composite laminate structure.
As a preferred improvement of the present invention, the stringer includes a stringer lower edge strip for securement to the skin, a stringer upper edge strip secured to the web, and a stringer web connecting the stringer upper edge strip and the stringer lower edge strip.
As a preferred improvement of the invention, the stringer upper edge strip is riveted to the web by a composite angle and rivet.
As a preferred improvement of the invention, the lower part of the bulkhead is provided with a through hole penetrating the web and the outer edge strip, the stringer passes through the through hole, and the stringer lower edge strip is abutted against the skin and fixed with the skin through rivets.
As a preferable improvement of the invention, the bulkhead divides the stringer into a left part and a right part which are broken, the stringer lower edge strip is abutted against the outer edge strip, and the stringer lower edge strip and the outer edge strip are fixed with the skin by rivets.
A method of designing an aircraft composite fuselage structure, the method comprising the steps of:
Step one, building a finite element model of a fuselage structure;
step two, endowing finite element model material properties;
Step three, meshing the airframe structure, and adopting a formula (1) to carry out hourglass control:
Wherein M is an overall mass matrix, C is an overall damping matrix, K is an overall stiffness matrix, Q (t) is an overall load vector, and x (t) is a displacement matrix;
step four, setting impact conditions;
step five, constructing a progressive damage model of the composite material to describe material failure, which specifically comprises the following steps:
assuming that the formers of the fuselage structure are orthotropic wire elastic materials, their damage constitutive relationship is represented by formulas (2) - (4):
Wherein 1 and 2 represent the longitudinal and transverse fiber directions, respectively; epsilon 11、ε22 and epsilon 12 represent strain vectors in the corresponding directions of the materials; σ 11、σ22 and σ 33 represent stress vectors in the corresponding directions of the materials; e 22 and E 11 each represent Young's modulus in the transverse and longitudinal directions; v 12 and v 21 generation of surface poisson ratio; g 33 represents the shear modulus; d 11、d22 and d 33 represent fiber damage variables in the machine direction, transverse direction, and in-plane shear, respectively;
a failure criterion is introduced to predict the failure of the finite element model, and the shear stress is coupled with the longitudinal and transverse positive stress of the material and is expressed by formulas (5) to (8):
fiber tensile failure:
fiber compression failure:
Matrix tensile failure:
matrix compression failure:
Wherein, F 1t、F1c、F2t and F 2c are respectively failure functions corresponding to longitudinal stretching, longitudinal compression, transverse stretching and transverse compression; x T and X C represent a longitudinal tensile strength and a longitudinal compressive strength, respectively; y T and Y C are transverse tensile strength and transverse compressive strength, respectively; s represents internal shear strength of the surface; And σ i (i= 11,22,12) represent effective stress and stress, respectively; alpha is a shearing effect weight, and the value range is 0 to 1;
introducing an exponential damage evolution criterion to reduce the rigidity of the material, and specifically comprises the following steps:
Longitudinal and lateral damage:
wherein d j (j=1t, 1c,2t,2 c) represents longitudinal and transverse damage variables, and d j includes longitudinal tensile failure, longitudinal compression failure, transverse tensile failure, and transverse compression failure; d j =0 means that no damage to the material has occurred, Representing complete damage to the material; l c represents a unit feature length; /(I)Representing the critical fracture energy of the material; /(I)Represents the elastic energy of the material; r j represents a damage threshold;
In-plane shear damage:
Wherein d 12 represents a shear damage variable, α 12 is determined by a material parameter; represents the maximum damage variable of shear damage, d 12 =0 represents that no damage occurs to the material,/> Representing complete damage to the material; r 12 represents a shear damage threshold;
step six, calculating by adopting an explicit finite element algorithm to obtain a simulation result, and adopting the following hourglass control method to reduce errors, wherein the method specifically comprises the following steps of:
Taking the initial time particle coordinate as X j (j=1, 2, 3), where at any time t the displacement of the particle X i (i=1, 2, 3) can be expressed by equation (11):
xi＝xi(Xj,t) (11)
the motion equation expressed by the matrix is obtained in consideration of the damping effect, and is expressed by formulas (12) to (15):
Fre＝Fex-Fin (13)
wherein M is an overall mass matrix, C is an overall damping matrix, x (t) is a displacement matrix, K is an overall stiffness matrix, Representing external force,/>Representing internal force,/>Representing residual forces;
The acceleration is calculated by utilizing the modified form of the motion differential equation, the node speed is calculated by utilizing the center differential, and the expression is expressed by a formula (16):
And then the node displacement is obtained by utilizing the node speed, and is expressed by a formula (17):
respectively outputting initial peak load, total kinetic energy of the fuselage structure and change of total internal energy along with time, specific energy absorption of each structural element and change of internal energy along with time in the process of falling collision of the fuselage structure;
step seven, comparing the failure modes and the energy absorption characteristics;
and step eight, finishing the final design.
As a preferred improvement of the invention, in step three, the lower region of the fuselage structure is divided by a mesh having a cell size of 2mm to 10mm, and the upper region of the bottom region of the fuselage structure is divided by a mesh having a cell size of 10mm to 20 mm.
As a preferred improvement of the invention, in the fourth step, the impact condition is that the fuselage structure is impacted vertically with a crash initial velocity symmetrical to the rigid ground, and the initial velocity is 7m/s-12m/s.
As a preferable improvement of the present invention, the comparative crashworthiness energy absorption characteristic index specifically includes:
the total kinetic energy attenuation speed of the machine body structure is required to be in a smooth descending trend, the attenuation speed is high, and the rapid descending is an unstable failure mode;
The sum of the bulkhead and the stringer specific energy absorption has a maximum value;
The sum of the internal energy of the formers occupies a greater ratio of the total internal energy of the structure;
in the whole crushing process, the highest initial peak value is in an initial stage, the initial peak load is relatively close to the average load value in the whole impact process, and the impact load is relatively stable.
The beneficial effects of the invention are as follows:
1. Compared with a typical frame, the frame provided by the invention adopts an integrally formed composite material laminated plate structure, has the characteristics of simple structure, easiness in manufacturing, light weight, good integrity and the like, and shows excellent mechanical properties when bearing impact load;
2. The stringers are arranged at the positions of the bulkhead frames in an interrupted manner, so that the integrity of the bulkhead frames is ensured, the structural performance of the bulkhead frames is greatly improved, the stringers are connected with the bulkhead frames by adopting the composite angle bars, the stability of the integral structure formed by the skin, the bulkhead frames and the stringers in bearing is ensured, and the bearing capacity of the fuselage structure is effectively improved;
3. According to the invention, the bulkhead is perforated, so that the arrangement of the stringers is facilitated, the continuity of the arrangement of the stringers is ensured, the bearing of the fuselage structure is positively influenced, and the stringers and the bulkhead are connected by adopting the composite angle material, so that the bearing capacity of the fuselage structure is greatly improved;
4. According to the failure criterion and the damage evolution criterion provided by the invention, the established finite element model can accurately capture the failure mode and the load response of the airframe structure under the action of impact load;
5. The invention provides a complete system fuselage structure design scheme, has strong maneuverability, and provides important support for the design of a composite fuselage structure;
6. the invention calculates the finite element model by adopting an explicit finite element algorithm, and provides various methods for controlling the occurrence of the hourglass phenomenon aiming at the possible hourglass phenomenon, so that the model can accurately capture the airframe structure response under impact load.
[ Description of the drawings ]
For a clearer description of the technical solutions of the embodiments of the present invention, the drawings that are needed in the description of the embodiments will be briefly introduced below, it being obvious that the drawings in the description below are only some embodiments of the present invention, and that other drawings can be obtained according to these drawings without inventive effort for a person skilled in the art, wherein:
FIG. 1 is a schematic view of a fuselage substructure;
FIG. 2 is a schematic cross-sectional view of a bulkhead according to the invention;
FIGS. 3 (a) to 3 (c) are schematic views of three different structures of a spacer provided by the present invention;
FIG. 4 is an enlarged view of the bulkhead in place of connection with the web;
FIG. 5 is a front view of a former without an aperturing process;
FIG. 6 is an enlarged view of a portion of the attachment of the non-perforated bulkhead to the stringer;
FIGS. 7 (a) and 7 (b) are schematic views of two former opening shapes provided by the present invention;
FIG. 8 is a front view of the perforated former;
FIG. 9 is an enlarged view of a portion of the attachment of the bulkhead to the stringer after opening the aperture;
FIG. 10 is a flow chart of a method of designing an aircraft composite fuselage structure according to the present invention;
In the figure:
1.A skin; 2. stringers; 3. a spacer frame; 4. an inner edge strip; 5. a web; 6. an outer edge strip; 7. i-shaped partition frame; 8. a Z-shaped bulkhead; 9. c-shaped partition frame; 10. a rivet; 11. a fuselage centerline; 12. a stringer lower edge strip; 13. stringer webs; 14. the stringer upper edge strip; 15. composite angle material; 16. square holes; 17. arc-shaped holes.
[ Detailed description ] of the invention
The technical solutions of the embodiments of the present invention will be clearly and completely described in the following in conjunction with the embodiments of the present invention, and it is obvious that the described embodiments are only some embodiments of the present invention, but not all embodiments. All other embodiments, which can be made by those skilled in the art based on the embodiments of the invention without making any inventive effort, are intended to be within the scope of the invention.
It should be noted that all directional indicators (such as up, down, left, right, front, and rear … …) in the embodiments of the present invention are merely used to explain the relative positional relationship, movement, etc. between the components in a particular posture (as shown in the drawings), and if the particular posture is changed, the directional indicator is changed accordingly.
Furthermore, descriptions such as those referred to as "first," "second," and the like, are provided for descriptive purposes only and are not to be construed as indicating or implying a relative importance or implying an order of magnitude of the indicated technical features in the present disclosure. Thus, a feature defining "a first" or "a second" may explicitly or implicitly include at least one such feature. In the description of the present invention, the meaning of "plurality" means at least two, for example, two, three, etc., unless specifically defined otherwise.
In the present invention, unless specifically stated and limited otherwise, the terms "connected," "affixed," and the like are to be construed broadly, and for example, "affixed" may be a fixed connection, a removable connection, or an integral body; can be mechanically or electrically connected; either directly or indirectly, through intermediaries, or both, may be in communication with each other or in interaction with each other, unless expressly defined otherwise. The specific meaning of the above terms in the present invention can be understood by those of ordinary skill in the art according to the specific circumstances.
In addition, the technical solutions of the embodiments of the present invention may be combined with each other, but it is necessary to be based on the fact that those skilled in the art can implement the technical solutions, and when the technical solutions are contradictory or cannot be implemented, the combination of the technical solutions should be considered as not existing, and not falling within the scope of protection claimed by the present invention.
Referring to fig. 1 and 2, an aircraft composite fuselage structure includes a skin 1, stringers 2 and a bulkhead 3, wherein the skin 1 is laid on the bulkhead 3, the stringers 2 and the bulkhead 3 are assembled to support the skin 1, the bulkhead 3 includes an outer edge strip 6 riveted with the skin 1, a web 5 bent and extended from the outer edge strip 6, and an inner edge strip 4 bent and extended from the web 5 and arranged parallel to the outer edge strip 6, the thicknesses t of the inner edge strip 4, the web 5 and the outer edge strip 6 are the same, and the inner length of the web 5 is longer than the lengths L1 and L2 of the outer edge strip 6 and the inner edge strip 4.
The inner edge strip 4 and the outer edge strip 6 are respectively perpendicular to the web 5, the inner edge strip 4 extends from the top of the web 5 to a single direction, the outer edge strip 6 simultaneously extends from the bottom of the web to an opposite direction, and the partition frame 3 adopts an integrally formed composite material laminated plate structure, and besides, the invention also provides three cross-sectional forms of an I-shaped partition frame 7, a Z-shaped partition frame 8 and a C-shaped partition frame 9, as shown in figures 3 (a), 3 (b) and 3 (C).
The crashworthiness of the fuselage structure is not only related to the cross-sectional form of the formers 3, stringers 2 etc., but the arrangement of the individual structures has a great influence on the crashworthiness. According to the fuselage structure provided by the invention, the bulkhead 3 is directly connected with the skin 1, the enlarged view of the connection position of the bulkhead 3 and the web 5 is shown in fig. 4, in order to reduce the quality of connecting pieces, the web 5 is taken as the center, and rivets 10 are respectively fixed on the left side and the right side of the outer edge strip 6 for riveting with the skin 1, so that the connection mode can meet the requirement of the connection quality.
The stringer 2 includes a stringer lower edge strip 12 for securing with the skin 1, a stringer upper edge strip 14 secured with the web 5, and a stringer web 13 connecting the stringer upper edge strip 14 and the stringer lower edge strip 12. The stringer top edge strip 14 is riveted to the web 5 by a composite angle 15 and rivets 10.
To facilitate the placement of stringers 2, the present invention proposes two treatments, open and non-open, to the bulkhead surface. To improve the structural integrity of the fuselage, the formers 3, stringers 2 and skin 1 are arranged appropriately, taking the example of "I" -stringers.
Two arrangements are proposed for the surface open pore and non-open pore of the bulkhead 3:
1) In order to ensure the structural properties of the former 3, no perforation is performed, and a front view of the former 3 is shown in fig. 5. The enlarged view of the connection part of the bulkhead 3 and the stringer 2 is shown in fig. 6, the stringer 2 is arranged in an interrupted manner at the position of the bulkhead 3, the stringer lower edge strip 12 is positioned above the external edge strip 6, the stringer web 13 is taken as the center, and a rivet 10 is respectively fixed on the left side and the right side of the stringer lower edge strip 12 to rivet the overlapped part of the stringer lower edge strip 12, the external edge strip 6 and the skin 1. The bulkhead 3 and the stringer 2 are connected by adopting a composite angle 15, and the concrete operation method is as follows: the composite angle 15 is riveted with the stringer upper edge strip 14 and the web 5 by two small-sized rivets 10 on the left and right, respectively, with the stringer web 13 as the center. The connecting method provided by the invention ensures the overall stability of the wallboard structure when bearing impact load.
2) In order to facilitate the integral arrangement of the stringers 2, the bulkhead 3 is perforated, in particular by the method of operation: the former web 5 and the former lower cap 6 are cut to form a through hole so that the stringer 2 just passes through the former 3 smoothly and the stringer lower cap 12 is in direct contact with the skin 1. The invention provides two opening forms of square holes 16 and arc holes 17 for the opening shape of the web 5, and the corners of the square holes are preferably overformed by adopting an arc shape, as shown in fig. 7 (a) and 7 (b). A front view of the perforated former 3 is shown in figure 8. The enlarged view of the connection of the former 3 to the stringer 2 after the opening is shown in figure 9, the stringer 2 is arranged continuously through the former 3, and the former 3 and the stringer 2 are respectively riveted directly to the skin 1 by rivets 10. The bulkhead 3 and the stringer 2 are connected by adopting a composite angle 15, and the concrete operation method is as follows: the composite angle 15 is riveted with the stringer upper edge strip 14 and the bulkhead web 5 by using left and right small-sized rivets 10 with the stringer web 13 as the center.
The fuselage bulkhead and the stringer provided by the invention adopt a composite material laminated plate structure, the manufacturing process is simple, the weight of the fuselage structure can be greatly reduced, and the crashworthiness of the fuselage structure is obviously improved. The invention provides two arrangement modes aiming at the opening and the non-opening of the bulkhead, so that the structural integrity of the fuselage is better, and the fuselage has enough stability when bearing impact load.
Referring again to fig. 10, the present invention also provides a method for designing an aircraft composite fuselage structure, the method comprising the steps of:
Step one, building a finite element model of a fuselage structure;
in order to more accurately simulate the failure process of the fuselage structure, a multipoint restraint unit or a beam unit is adopted to simulate a riveting mode; taking actual drop collision conditions into consideration, adopting a rigid surface to simulate rigid ground; in consideration of the long time spent by adopting the high-order shell units, the high-order shell units are very sensitive to grid division, and the calculation accuracy is greatly reduced due to unit distortion, the invention provides the modeling of the fuselage skin, the bulkhead and the stringer by adopting the four-node shell units, and the finite element model is consistent with the actual fuselage structure size by adjusting the thickness of the shell units. Specifically, the fuselage structure includes skins, formers and stringers, and the geometric dimensions of the skins, formers and stringers may be sized according to the geometric dimensions of the skins, formers and stringers of a typical fuselage structure.
Step two, endowing finite element model material properties;
It should be further noted that the finite element models of fuselage structures employing formers and stringers of different cross-sectional forms are given multiple sets of different material properties. The skin is mainly used for maintaining the appearance of the fuselage, and is made of epoxy carbon fiber composite materials with excellent performance characteristics such as high temperature resistance, friction resistance, corrosion resistance and the like, and the bulkhead and the stringer are made of resin matrix composite materials or metal/carbon fiber mixed materials with excellent mechanical properties such as good corrosion resistance, high buckling strength and the like. The rivet material adopts aluminum alloy or stainless steel or titanium alloy material, etc.
Step three, meshing the airframe structure, and adopting a formula (1) to carry out hourglass control:
Wherein M is an overall mass matrix, C is an overall damping matrix, K is an overall stiffness matrix, Q (t) is an overall load vector, and x (t) is a displacement matrix.
It should be further noted that, in order to ensure the calculation accuracy, the structure is meshed by adopting a quadrilateral mesh as much as possible. The grid division is to show the geometric characteristics of the structure, and the key parts of the falling collision such as the joint, the bottom of the machine body and the like adopt high-quality matrix units with the length-width ratio close to 1 and refine the grids; smaller grid sizes result in less computationally efficient, while excessive grid sizes result in greater overall error. The bottom of the fuselage structure is a main energy absorption area in the process of falling and bumping, the structural deformation is very obvious, in order to ensure the calculation accuracy, a denser grid is adopted, therefore, the lower area of the fuselage structure is divided by adopting grids with the unit size of 2mm-10mm, and the upper area of the bottom area of the fuselage structure is divided by adopting grids with the unit size of 10mm-20 mm.
Of course, the cell density can be adjusted as needed for analysis. The adoption of the reduction integration method can cause the quadrilateral shell unit to generate a zero-energy mode without energy consumption or an hourglass mode, and the hourglass phenomenon can cause larger errors between an analysis result and an actual result. In the aspect of grid division, the quality of grids needs to be improved, and the grids are properly thinned to avoid sharp points, because the sharp points are easy to generate hourglass under the action of concentrated force in the collision process; in order to prevent the occurrence of the hourglass phenomenon, the volume viscosity coefficient of the structure can be properly improved or a full-order integral unit formula can be used; in order to inhibit the zero energy mode, the invention considers the damping effect in the motion equation, and adds an hourglass damping term, so that the hourglass damping method can well solve the hourglass phenomenon, and a more accurate simulation result is obtained, and the equation is shown in a formula (1).
Step four, setting impact conditions;
Specifically, the invention mainly considers the vertical impact condition. The invention applies fixed constraint to the rigid plane of the simulated ground, and applies a symmetrical vertical impact rigid ground with the initial impact speed to the machine body structure model. To reduce the strain rate sensitivity of the material, the initial velocity may be set to 7m/s to 12m/s, and the influence of gravitational acceleration is taken into consideration.
Step five, constructing a progressive damage model of the composite material to describe material failure;
It should be noted that the model predicts material failure based on macroscopic scale using a failure criterion that considers multiple failure modes, taking into account the effects of shear stress in the transverse and longitudinal failure of the material. When the damage of the material meets the initial failure criterion, a damage variable is introduced to describe the damage evolution process of the material, and a damage evolution criterion is adopted to predict the damage expansion of the material. A large number of simulation experiments show that the damage model can accurately capture the failure mode and the energy absorption characteristic of the material, and specifically comprises the following steps:
assuming that the formers of the fuselage structure are orthotropic wire elastic materials, their damage constitutive relationship is represented by formulas (2) - (4):
Wherein 1 and 2 represent the longitudinal and transverse fiber directions, respectively; epsilon 11、ε22 and epsilon 12 represent strain vectors in the corresponding directions of the materials; σ 11、σ22 and σ 33 represent stress vectors in the corresponding directions of the materials; e 22 and E 11 each represent Young's modulus in the transverse and longitudinal directions; v 12 and v 21 generation of surface poisson ratio; g 33 represents the shear modulus; d 11、d22 and d 33 represent fiber damage variables in the machine direction, transverse direction, and in-plane shear, respectively;
a failure criterion is introduced to predict the failure of the finite element model, and the shear stress is coupled with the longitudinal and transverse positive stress of the material and is expressed by formulas (5) to (8):
fiber tensile failure:
fiber compression failure:
Matrix tensile failure:
matrix compression failure:
Wherein, F 1t、F1c、F2t and F 2c are respectively failure functions corresponding to longitudinal stretching, longitudinal compression, transverse stretching and transverse compression; x T and X C represent a longitudinal tensile strength and a longitudinal compressive strength, respectively; y T and Y C are transverse tensile strength and transverse compressive strength, respectively; s represents internal shear strength of the surface; And σ i (i= 11,22,12) represent effective stress and stress, respectively; alpha is a shearing effect weight, and the value range is 0 to 1;
introducing an exponential damage evolution criterion to reduce the rigidity of the material, and specifically comprises the following steps:
Longitudinal and lateral damage:
wherein d j (j=1t, 1c,2t,2 c) represents longitudinal and transverse damage variables, and d j includes longitudinal tensile failure, longitudinal compression failure, transverse tensile failure, and transverse compression failure; d j =0 means that no damage to the material has occurred, Representing complete damage to the material; l c represents a unit feature length; /(I)Representing the critical fracture energy of the material; /(I)Represents the elastic energy of the material; r j represents a damage threshold;
In-plane shear damage:
Wherein d 12 represents a shear damage variable, α 12 is determined by a material parameter; represents the maximum damage variable of shear damage, d 12 =0 represents that no damage occurs to the material,/> Representing complete damage to the material;
step six, calculating by adopting an explicit finite element algorithm to obtain a simulation result, and adopting the following hourglass control method to reduce errors;
It should be noted that, the explicit algorithm adopts the center difference in the dynamic equation to solve, and does not need balanced iteration, so that the method has higher calculation efficiency, and is superior to the implicit algorithm when solving the impact problem. The invention adopts an explicit finite element algorithm to calculate the finite element model, but the method for reducing integral adopted in the explicit algorithm is easy to cause an hourglass, and the invention adopts the following hourglass control method to reduce errors, and specifically comprises the following steps:
Taking the initial time particle coordinate as X j (j=1, 2, 3), where at any time t the displacement of the particle X i (i=1, 2, 3) can be expressed by equation (11):
xi＝xi(Xj,t) (11)
the motion equation expressed by the matrix is obtained in consideration of the damping effect, and is expressed by formulas (12) to (15):
Fre＝Fex-Fin (13)
Wherein M is an overall mass matrix, C is an overall damping matrix, x (t) is a displacement matrix, K is an overall stiffness matrix, F n ex represents an external force, Representing internal force,/>Representing residual forces;
The acceleration is calculated by utilizing the modified form of the motion differential equation, the node speed is calculated by utilizing the center differential, and the expression is expressed by a formula (16):
And then the node displacement is obtained by utilizing the node speed, and is expressed by a formula (17):
respectively outputting initial peak load, total kinetic energy of the fuselage structure and change of total internal energy along with time, specific energy absorption of each structural element and change of internal energy along with time in the process of falling collision of the fuselage structure;
it should be further noted that specific energy absorption is defined as the amount of energy absorbed per unit mass of the fuselage structure.
Step seven, comparing the failure modes and the energy absorption characteristics;
specifically comprises the steps of comparing the initial failure state of the structural element and the impact energy absorption characteristic index.
Wherein, comparing the initial failure modes of the structural elements specifically comprises comparing failure modes of the lower structure of the fuselage when 100ms-250ms is impacted or failure modes when the crushing displacement is 50% of the maximum displacement. The fuselage structure failure mode with excellent energy absorption performance should have the following characteristics: the bottom frame of the machine body is not broken, and the structural elements are not subjected to unstable failure modes such as large-scale breaking, large-scale buckling and the like; the structural deformation of the whole fuselage is not concentrated at the bottom of the fuselage, the gradual failure process is followed, the bottom in the fuselage is deformed firstly, and the deformation is expanded to the two ends of the fuselage along with the continuous progress of the impact process; with sufficient effective energy absorbing area.
Wherein, the comparative crashworthiness energy absorption characteristic index specifically comprises: the total kinetic energy attenuation speed of the machine body structure is required to be in a smooth descending trend, the attenuation speed is high, and the rapid descending is an unstable failure mode; the sum of the bulkhead and the stringer specific energy absorption has a maximum value; the sum of the internal energy of the formers occupies a greater ratio of the total internal energy of the structure; in the whole crushing process, the highest initial peak value is in an initial stage, the initial peak load is relatively close to the average load value in the whole impact process, and the impact load is relatively stable.
And step eight, finishing the final design.
In particular, the final design includes the optimal material parameters and geometry of the fuselage structure.
The beneficial effects of the invention are as follows:
1. Compared with a typical frame, the frame provided by the invention adopts an integrally formed composite material laminated plate structure, has the characteristics of simple structure, easiness in manufacturing, light weight, good integrity and the like, and shows excellent mechanical properties when bearing impact load;
2. The stringers are arranged at the positions of the bulkhead frames in an interrupted manner, so that the integrity of the bulkhead frames is ensured, the structural performance of the bulkhead frames is greatly improved, the stringers are connected with the bulkhead frames by adopting the composite angle bars, the stability of the integral structure formed by the skin, the bulkhead frames and the stringers in bearing is ensured, and the bearing capacity of the fuselage structure is effectively improved;
3. According to the invention, the bulkhead is perforated, so that the arrangement of the stringers is facilitated, the continuity of the arrangement of the stringers is ensured, the bearing of the fuselage structure is positively influenced, and the stringers and the bulkhead are connected by adopting the composite angle material, so that the bearing capacity of the fuselage structure is greatly improved;
4. According to the failure criterion and the damage evolution criterion provided by the invention, the established finite element model can accurately capture the failure mode and the load response of the airframe structure under the action of impact load;
5. The invention provides a complete system fuselage structure design scheme, has strong maneuverability, and provides important support for the design of a composite fuselage structure;
6. the invention calculates the finite element model by adopting an explicit finite element algorithm, and provides various methods for controlling the occurrence of the hourglass phenomenon aiming at the possible hourglass phenomenon, so that the model can accurately capture the airframe structure response under impact load.
Although embodiments of the present invention have been disclosed above, it is not limited to the details and embodiments shown and described, it is well suited to various fields of use for which the invention would be readily apparent to those skilled in the art, and accordingly, the invention is not limited to the specific details and illustrations shown and described herein, without departing from the general concepts defined in the claims and their equivalents.
Claims (8)
1. A method of designing an aircraft composite fuselage structure, the method comprising the steps of:
Step one, building a finite element model of a fuselage structure, wherein the fuselage structure comprises a skin, stringers and a bulkhead, the skin is laid on the bulkhead, the stringers and the bulkhead are assembled to support the skin, the bulkhead adopts an integrally formed composite material laminated plate structure, the bulkhead comprises an outer edge strip riveted with the skin, a web plate bent and extended from the outer edge strip and an inner edge strip bent and extended from the web plate and arranged in parallel with the outer edge strip, the thickness of the inner edge strip, the web plate and the outer edge strip is the same, and the inner side length of the web plate is longer than the lengths of the outer edge strip and the inner edge strip;
step two, endowing finite element model material properties;
Step three, meshing the airframe structure, and adopting a formula (1) to carry out hourglass control:
wherein M is an overall mass matrix, C is an overall damping matrix, K is an overall stiffness matrix, and Q (t) is an overall load vector; x (t) is a displacement matrix;
step four, setting impact conditions;
step five, constructing a progressive damage model of the composite material to describe material failure, which specifically comprises the following steps:
assuming that the formers of the fuselage structure are orthotropic wire elastic materials, their damage constitutive relationship is represented by formulas (2) - (4):
Wherein 1 and 2 represent the longitudinal and transverse fiber directions, respectively; epsilon 11、ε22 and epsilon 12 represent strain vectors in the corresponding directions of the materials; σ 11、σ22 and σ 33 represent stress vectors in the corresponding directions of the materials; e 22 and E 11 each represent Young's modulus in the transverse and longitudinal directions; v 12 and v 21 generation of surface poisson ratio; g 33 represents the shear modulus; d 11、d22 and d 33 represent fiber damage variables in the machine direction, transverse direction, and in-plane shear, respectively;
a failure criterion is introduced to predict the failure of the finite element model, and the shear stress is coupled with the longitudinal and transverse positive stress of the material and is expressed by formulas (5) to (8):
fiber tensile failure:
fiber compression failure:
Matrix tensile failure:
matrix compression failure:
Wherein, F 1t、F1c、F2t and F 2c are respectively failure functions corresponding to longitudinal stretching, longitudinal compression, transverse stretching and transverse compression; x T and X C represent a longitudinal tensile strength and a longitudinal compressive strength, respectively; y T and Y C are transverse tensile strength and transverse compressive strength, respectively; s represents internal shear strength of the surface; And σ i (i= 11,22,12) represent effective stress and stress, respectively; alpha is a shearing effect weight, and the value range is 0 to 1;
introducing an exponential damage evolution criterion to reduce the rigidity of the material, and specifically comprises the following steps:
Longitudinal and lateral damage:
wherein d j (j=1t, 1c,2t,2 c) represents longitudinal and transverse damage variables, and d j includes longitudinal tensile failure, longitudinal compression failure, transverse tensile failure, and transverse compression failure; d j =0 means that no damage to the material has occurred, Representing complete damage to the material; l c represents a unit feature length; /(I)Representing the critical fracture energy of the material; /(I)Represents the elastic energy of the material; r j represents a damage threshold;
In-plane shear damage:
Wherein d 12 represents a shear damage variable, α 12 is determined by a material parameter; represents the maximum damage variable of shear damage, d 12 =0 represents that no damage occurs to the material,/> Representing complete damage to the material; r 12 represents a shear damage threshold;
step six, calculating by adopting an explicit finite element algorithm to obtain a simulation result, and adopting the following hourglass control method to reduce errors, wherein the method specifically comprises the following steps of:
Taking the initial time particle coordinate as X j (j=1, 2, 3), where at any time t the displacement of the particle X i (i=1, 2, 3) can be expressed by equation (11):
xi＝xi(Xj,t) (11)
the motion equation expressed by the matrix is obtained in consideration of the damping effect, and is expressed by formulas (12) to (15):
Fre＝Fex-Fin (13)
wherein M is an overall mass matrix, C is an overall damping matrix, x (t) is a displacement matrix, K is an overall stiffness matrix, Representing external force,/>Representing internal force,/>Representing residual forces;
The acceleration is calculated by utilizing the modified form of the motion differential equation, the node speed is calculated by utilizing the center differential, and the expression is expressed by a formula (16):
And then the node displacement is obtained by utilizing the node speed, and is expressed by a formula (17):
respectively outputting initial peak load, total kinetic energy of the fuselage structure and change of total internal energy along with time, specific energy absorption of each structural element and change of internal energy along with time in the process of falling collision of the fuselage structure;
step seven, comparing the failure modes and the energy absorption characteristics;
and step eight, finishing the final design.
2. A method of designing an aircraft composite fuselage structure according to claim 1, wherein: in step one, the stringer includes a stringer lower cap for securing to a skin, a stringer upper cap secured to a web, and a stringer web connecting the stringer upper cap and the stringer lower cap.
3. A method of designing an aircraft composite fuselage structure according to claim 2, wherein: in the first step, the stringer upper edge strip is riveted and fixed with the web by a composite angle and a rivet.
4. A method of designing an aircraft composite fuselage structure according to claim 2, wherein: in the first step, a through hole penetrating through the web plate and the outer edge strip is formed in the lower portion of the bulkhead, the stringer penetrates through the through hole, and the stringer lower edge strip is abutted to the skin and fixed with the skin through rivets.
5. A method of designing an aircraft composite fuselage structure according to claim 2, wherein: in the first step, the bulkhead divides the stringer into a left part and a right part which are disconnected, the stringer lower edge strip is abutted on the outer edge strip, and the stringer lower edge strip and the outer edge strip are fixed with the skin through rivets.
6. A method of designing an aircraft composite fuselage structure according to claim 1, wherein: in the third step, the lower area of the airframe structure is divided by adopting grids with the unit size of 2mm-10mm, and the upper area of the bottom area of the airframe structure is divided by adopting grids with the unit size of 10mm-20 mm.
7. A method of designing an aircraft composite fuselage structure according to claim 1, wherein: in the fourth step, the impact condition is that the fuselage structure is impacted vertically by applying a crash initial speed which is symmetrical to the rigid ground, and the initial speed is 7m/s-12m/s.
8. A method of designing an aircraft composite fuselage structure according to claim 1, wherein: the comparative crashworthiness energy absorption characteristic index specifically comprises:
the total kinetic energy attenuation speed of the machine body structure is required to be in a smooth descending trend, the attenuation speed is high, and the rapid descending is an unstable failure mode;
The sum of the bulkhead and the stringer specific energy absorption has a maximum value;
The sum of the internal energy of the formers occupies a greater ratio of the total internal energy of the structure;
in the whole crushing process, the highest initial peak value is in an initial stage, the initial peak load is relatively close to the average load value in the whole impact process, and the impact load is relatively stable.
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