MSC.Nastran analysis solution-based general aviation aircraft composite spar structure simulation method

Abstract

The invention belongs to the technical field of aircraft composite material structure design and analysis, and provides a general aviation aircraft composite material spar structure simulation method based on MSC.Nastran analysis solution, and the method comprises the steps: building a wing overall finite element model; carrying out attribute definition on the wing model structure to obtain correct attributes of a rod CROD unit, endowing the correct attributes into a finite element model, and establishing the finite element model; load working condition calculation is applied to the established finite element model; extracting a rod unit force, and correcting a rod CROD unit load by using the correction factor to obtain an actual load; according to the section rigidity proportion, the actual load is distributed to a skin boss and a beam edge strip, and respective strength checking is carried out. According to the method, the effect of the spar structure in the aircraft can be accurately simulated, the structural design and strength analysis efficiency of the aircraft can be improved, the development period of the composite spar of the general aviation aircraft is shortened, and the method is suitable for analysis and design iteration of the composite spar of the general aviation aircraft.

Description

Technical Field

[0001] The present invention relates to the technical field of aircraft composite material structure design and analysis, in particular to a method for simulating composite wing spar structures of general aviation aircraft based on MSC.Nastran analysis and solution. Background Art

[0002] In recent years, with the country's attention and development of the general aviation industry, many general aviation aircraft models have gradually emerged in my country, from the Eagle 500, the first domestically produced light aircraft designed, tested and produced in full compliance with the CCAR-23 aviation regulations, to the current Ruixiang series, Aurora series, Tongfei AG series and GA20 and many other models. In addition, many foreign mature models have also been introduced, such as the SR series of Cirrus Aircraft Corporation of the United States and the DA series of aircraft of Diamond Aircraft Corporation of Austria. They have entered the Chinese market in various forms. These aircraft have passed the corresponding 23 airworthiness certifications and obtained airworthiness certificates. The vigorous development of the general aviation industry has brought more opportunities to general aviation aircraft model designers, but also brought challenges. How to conduct aircraft structure simulation conveniently and accurately has become the focus of designers and airworthiness certification.

[0003] The use of composite materials is a key topic in the design and manufacture of today's aircraft models. Composites can even be used to symbolize and represent an aircraft's low economy and superior performance. The trend toward composite materials in aircraft structures is becoming increasingly clear with the continuous approval of new models. Composite materials are also increasingly being used in the design and manufacture of general aviation aircraft structures. This is because economic efficiency is crucial to the general aviation industry, and economic efficiency is primarily determined by two factors: structural manufacturing cost and subsequent maintenance and repair costs. For example, Diamond Aircraft's DA40 aircraft maintenance manual clearly states that regular inspections of major structures are an integral part of the aircraft's indefinite lifespan. The DA40 utilizes extensive composite materials to enhance economic efficiency, and the aforementioned major structures are also manufactured from composite materials, particularly the wing spars, fuselage frames, and skins, which carry the most load. The wing spars are the most critical structures for general aviation aircraft, bearing the bending moments and shear forces generated by the enormous aerodynamic loads. Therefore, they require sufficient strength and stiffness, making the simulation methods used in their design and analysis crucial.

[0004] Currently, the simulation and analysis of composite wing spars for general aviation aircraft mostly uses the detailed finite element method (DFE), which simulates the wing spar cap and skin together. This method is labor-intensive, complex, and lacks versatility. Small structural changes can lead to significant changes in the model. Furthermore, it suffers from low computational efficiency, difficult iterations, and inability to account for all failure modes of the wing spar structure. For these reasons, there is an urgent need to develop an applicable and convenient method for simulating composite wing spars for general aviation aircraft. Summary of the Invention

[0005] In view of this, the present invention discloses a simulation method for a composite wing spar structure of a general aviation aircraft based on MSC.Nastran analysis and solution, so as to accurately simulate the wing spar stiffness of the composite material through a coarse model.

[0006] The technical solution provided by the present invention is a method for simulating composite wing spar structures of general aviation aircraft based on MSC.Nastran analysis and solution, including:

[0007] S1: Establish the finite element model of the entire wing;

[0008] S2: Define the properties of the wing model structure, obtain the correct properties of the rod CROD unit, assign the correct properties to the finite element model, and establish the finite element model;

[0009] When calculating the structural properties of the spar, each half of the spar cross section is divided into three regions: Region 1: the skin at the spar flange, Region 2: the skin boss and spar flange at the spar flange, and Region 3: the spar web. The skin boss and spar flange in Region 2 are constructed by layup and simulated using rod CROD elements.

[0010] S3: Apply load calculation to the established finite element model; extract the rod unit force in region 2 from the calculation results, and use the correction factor to correct the rod CROD unit load to obtain the actual load; the correction factor is h fem / h real , where h real h represents the distance between the neutral axis of the beam and the neutral axis of zone 2; fem represents the distance between the neutral axis of the beam and the position of the bar element;

[0011] S4: Based on the cross-sectional stiffness ratio, the actual load is distributed to the skin boss and beam flange, and their respective strength checks are performed.

[0012] Preferably, the step S1 of establishing the overall finite element model of the wing includes:

[0013] S11: extracting the outer surface of the skin from the digital model of the general aviation aircraft wing structure as a reference for constructing the finite element model, and adjusting the normal of the outer surface to point to the inside of the box segment;

[0014] S12: Based on the benchmark, establish a finite element model of the composite wing spar and other structures of the general aviation aircraft;

[0015] The Hypermesh software is used to establish a finite element model of the composite wing spar and other structures of a general aviation aircraft. The other structures described in S12 include wing spar strips, reinforcements, and wing spar webs. The wing spar strips and reinforcements are simulated using rod Crod units, and the wing spar webs are simulated using shell units Cquad4 and Ctria3.

[0016] Specifically, in addition to the wing spar structure, S2 defines the properties of other wing model structures, including: the skin, ribs, and beam web structures use the Pcomp paving method to define the properties and lay them out according to the actual situation of the structural design.

[0017] Specifically, in S2, based on the principle of consistency of moment of inertia, the structural properties of the wing spar are calculated, including;

[0018] S21: Calculate the distance between the neutral axis shared by the upper and lower regions 2 of the spar and the reference position as shown in formula (1):

[0019]

[0020] Where: A i — Areas of upper / lower skin bosses and upper / lower spar caps;

[0021] E i ——Equivalent elastic modulus of the upper / lower skin bosses and upper / lower spar caps;

[0022] c i — The distance between the upper / lower skin boss and the neutral axis of the upper / lower spar cap relative to the reference position;

[0023] i=1 is the upper skin boss; i=2 is the upper edge of the beam; i=3 is the lower edge of the beam; i=4 is the lower skin

[0024] The formula for calculating the distance from the neutral axis of the upper and lower regions 2 of the spar to the reference position is:

[0025]

[0026] Where: A i — Areas of upper / lower skin bosses and upper / lower spar caps;

[0027] E i ——Equivalent elastic modulus of the upper / lower skin bosses and upper / lower spar caps;

[0028] c i — The distance between the upper / lower skin boss and the neutral axis of the upper / lower spar cap relative to the reference position;

[0029] i=1 is the upper / lower skin boss; i=2 is the upper / lower spar cap;

[0030] S23: Calculate the distance between the neutral axis of the upper and lower regions 2 and the shared neutral axis. The calculation formula is as follows:

[0031]

[0032] Where: hreal_upper ——The distance between the neutral axis of upper region 2 and the shared neutral axis;

[0033] h real_lower - the distance between the neutral axis of lower region 2 and the shared neutral axis;

[0034] C g - the distance between the neutral axis shared by the upper and lower parts of area 2 and the reference position;

[0035] C real_upper - the distance from the neutral axis of the upper area 2 itself to the reference position;

[0036] C real_lower - the distance from the neutral axis of the lower area 2 itself to the reference position;

[0037] S24: According to equations (1)-(3), the area of ​​the rod CROD unit introduced in the finite element model is:

[0038]

[0039] Where: E fem ——The equivalent elastic modulus assigned to region 2 in the model;

[0040] h fem —the distance between the neutral axis of the beam and the position of the rod element;

[0041] h real — the distance between the neutral axis of the beam and the neutral axis of zone 2;

[0042] E i_real

[0043] A i — Areas of upper / lower skin bosses and upper / lower spar caps;

[0044] S25: Modify the properties of the beam flange and skin boss. The modification principle includes: assuming that the moment of inertia of the finite element unit is equal to that of the actual structure, and simulating the action and load of the wing beam according to formula (5):

[0045]

[0046] Where: h real — the distance between the neutral axis of the beam and the neutral axis of zone 2;

[0047] h fem —the distance between the neutral axis of the beam and the position of the rod element;

[0048] A real ——The actual area of ​​Region 2;

[0049] Afem —The area assigned to Region 2 in the model;

[0050] E real ——The true equivalent elastic modulus of region 2;

[0051] E fem ——The equivalent elastic modulus assigned to region 2 in the model.

[0052] Specifically, in S3, MSC.Nastran software is used to apply load conditions to the finite element model.

[0053] The present invention provides a method for simulating the composite wing spar structure of a general aviation aircraft based on MSC.Nastran analysis and solution. The method utilizes the commercial finite element analysis software MSC.Nastran to perform analysis and obtain internal force solutions. Specifically, a plate-rod system is used to simulate the wing spar of the general aviation aircraft. Based on the principle of consistent moments of inertia, the correct properties of the rod system are calculated and assigned to the finite element model. A true load condition is applied to the model and an internal force solution analysis is performed. The internal force solution proposed by the spar cap combination is corrected to obtain the true load. The load is distributed to the spar cap and the skin boss according to the stiffness ratio. The spar is verified and analyzed using an engineering analysis method, thereby providing accurate input for the analysis and iterative design of the composite wing spar of the general aviation aircraft.

[0054] This method can fully consider the structural characteristics and accurately simulate the role of the wing spar structure in the aircraft. The method of the present invention also has the characteristics of convenient modeling, high calculation efficiency, and accurate load transmission, which can improve the efficiency of aircraft structural design and strength analysis. In addition, combined with engineering analysis methods or tools, it can achieve rapid analysis and iterative design of the wing spar structure with the most severe loads for general aviation aircraft, shorten the development cycle of composite wing spars for general aviation aircraft, enter the verification stage as soon as possible, and form a set of analysis and design iterations suitable for composite wing spars for general aviation aircraft.

[0055] It is to be understood that the foregoing general description and the following detailed description are exemplary and explanatory only and are not restrictive of the disclosure of the present invention. BRIEF DESCRIPTION OF THE DRAWINGS

[0056] The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments consistent with the invention and, together with the description, serve to explain the principles of the invention.

[0057] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the following briefly introduces the drawings required for use in the embodiments or the description of the prior art. Obviously, for ordinary technicians in this field, other drawings can be obtained based on these drawings without any creative work.

[0058] Figure 1A flow chart of a method for simulating a composite wing spar structure of a general aviation aircraft based on MSC.Nastran analysis and solution according to an embodiment of the present invention;

[0059] Figure 2 A finite element model of the entire wing provided for the disclosed embodiment of the present invention;

[0060] Figure 3 A cross-sectional view of an I-shaped wing spar provided in accordance with an embodiment of the present invention;

[0061] Figure 4 A finite element model of a wing provided in accordance with an embodiment of the present invention;

[0062] Figure 5 The ply properties of a certain portion of the skin provided by the disclosed embodiments of the present invention;

[0063] Figure 6 A diagram showing the division of the upper half of a beam according to an embodiment of the present invention;

[0064] Figure 7 Schematic diagram of the distance used in beam property calculation provided in the embodiment disclosed in the present invention. DETAILED DESCRIPTION

[0065] Exemplary embodiments will now be described in detail, examples of which are illustrated in the accompanying drawings. In the following description, when referring to the drawings, like numbers in different figures represent the same or similar elements unless otherwise indicated. The embodiments described in the following exemplary embodiments are not intended to represent all possible embodiments consistent with the present invention. Rather, they are merely examples of systems consistent with certain aspects of the present invention, as detailed in the appended claims.

[0066] Prior art has many deficiencies in wing spar simulation and analytical calculations for composite materials used in general aviation aircraft. For example, simulation of composite wing spar structures for general aviation aircraft can only be performed through detailed models, complex modeling, and inefficient calculation methods. Therefore, the present invention provides a method for simulating composite wing spar structures for general aviation aircraft based on MSC.Nastran analysis and solution. This method accurately simulates the stiffness of composite wing spar using a coarse model, and features convenient modeling, high computational efficiency, and accurate load transfer. Combined with analytical tools, this method enables rapid analysis and iterative design of wing spar structures for general aviation aircraft under the most severe loads.

[0067] like Figure 1 Shown is a schematic flow diagram of the method of the present invention, which specifically includes:

[0068] S1: Establish the finite element model of the entire wing;

[0069] Extract the outer surface of the skin from the digital model of the general aviation aircraft wing structure as the basis for constructing the finite element model, and adjust the normal of the surface to point to the inside of the box segment;

[0070] The finite element model of the composite wing spar and other structures of a general aviation aircraft was established using Hypermesh software. It should be noted that the spar caps and reinforcements are simulated using rod elements (Crod), and the spar webs are simulated using shell elements with bending capabilities (Cquad4 and Ctria3). The finite element model of the entire wing is attached. Figure 4 .

[0071] S2: Define the properties of the wing model structure:

[0072] The skin, ribs, beam webs and other structures are given attributes by Pcomp paving method, and the layers are laid according to the actual situation of the structural design, such as Figure 5 The layer properties of a certain part of the skin.

[0073] The spar cap properties are calculated and determined in the following way;

[0074] 1) Each half of the beam section is divided into three regions: Region 1, the skin at the beam flange, Region 2, the skin boss at the beam flange and the beam flange, and Region 3, the beam web. Region 2 of the spar is composed of the skin boss and the beam flange, and can be considered as consisting of two parts of the layup stacked together during modeling. In the finite element (FEM) model, this stacking is simulated using rod elements (CROD), and the attached Figure 6 Divide the diagram for the upper half of the beam;

[0075] 2) Calculate the neutral axis shared by the upper and lower regions of the spar, see the attached Figure 7 , the distance between the neutral axis and the reference position is calculated as follows:

[0076]

[0077] Where: A i --area;

[0078] E i ——Equivalent elastic modulus of laminate;

[0079] c i —the distance of the neutral axis of each part relative to the reference position;

[0080] i=1 is the upper skin boss; i=2 is the upper edge strip of the beam; i=3 is the lower edge strip of the beam; i=4 is the lower skin boss.

[0081] Attachment Figure 7 The distance calculation formula for each part in is as follows, where t s_upper and t s_lower are the thickness of the upper and lower skin respectively.

[0082]

[0083] The spar at rib 1 (representing one of many ribs) is selected as an example for calculation to determine the common neutral axis of the spar here, see Table 1 below.

[0084]

[0085] 3) Based on the above segmentation of the beam, the rod (CROD) unit is formed by sequentially paving the beam flange and the skin boss. The distance formula from the neutral axis of the paving structure itself to the reference position is:

[0086]

[0087] Where: i=1 is the upper / lower skin boss; i=2 is the upper / lower edge strip of the beam.

[0088] 4) Calculate the distance from the neutral axis of the upper and lower regions 2 to the common neutral axis using the following formula:

[0089] h real_upper =C real_upper -C g

[0090] h real_lower =C g -C real_lower

[0091] Where: h real_upper - the distance between the upper region 2 and the common neutral axis;

[0092] h real_lower - the distance between the lower region 2 and the common neutral axis;

[0093] C g — the distance between the common neutral axis and the reference position;

[0094] C real_upper - the distance between upper area 2 and the reference position;

[0095] C real_lower - the distance between lower area 2 and the reference position;

[0096] Then, using the first two rows of data in Table 1, we can get C real_upper = 822 mm, then the actual distance from the upper area simulated by the rod element to the neutral axis is h real_upper =C real_upper -C g =822-527=295mm.

[0097] Using the last two rows of data in the table above, we can find C real_upper =11mm, then the actual distance from the lower area simulated by the rod element to the neutral axis is

[0098] 5) According to the formulas in steps 2)-4), the area of ​​the rod element introduced in the finite element is:

[0099]

[0100] Where: E fem ——Equivalent elastic modulus of the beam flange and skin boss after laying in sequence;

[0101] h fem ——The location of the CROD unit in the model;

[0102] Still taking the above beam section as an example, see the attached Figure 7 , the distance between the upper and lower nodes is H = 838mm. According to the above formula, the area of ​​the rod element introduced in the finite element is:

[0103]

[0104] Where: h fem_upper =HC g =311mm;

[0105] h fem_lower =C g =527mm;

[0106] 6) As attached Figure 7 As shown, the CROD is located at a node in the FEM, and its distance from the neutral axis differs from the actual distance of beam region 2. Therefore, the properties of the beam cap (CAP) and skin pad must be modified before they can be included in the FEM. The principle of modification is to assume that the finite element moment of inertia is equal to that of the actual structure. This allows for a realistic simulation of the cap's function and load-bearing conditions. The formula is as follows:

[0107]

[0108] Where: h real — the distance between the neutral axis of the beam and the neutral axis of zone 2;

[0109] h fem —the distance between the neutral axis of the beam and the position of the rod element;

[0110] A real ——the actual area of ​​beam region 2;

[0111] A fem ——The area assigned to beam region 2 in the model;

[0112] E real ——real equivalent modulus of beam region 2;

[0113] E fem ——Equivalent modulus of beam region 2 assigned in the model;

[0114] S3: Apply load calculation to the established finite element model;

[0115] After assigning attributes to the model, apply load conditions and perform calculations using MSC.Nastran software. The unit forces are extracted from the calculation results. The CROD unit loads must be corrected with a correction factor of h. fem / h real , the correction factors of the upper and lower rod elements in the above example are 311 / 295=1.05 and 527 / 516=1.02 respectively.

[0116] S4: Analyze and check the structure;

[0117] Extract the Crod element loads in simulation area 2, modify the loads using the load correction factor from step 5, and redistribute the loads between the beam flange and the skin boss according to the stiffness ratio for analysis and verification.

[0118] The above-established simulation method for the composite wing spar structure of general aviation aircraft allows for convenient model establishment; the plate rod system can accurately simulate the load-bearing, load-transferring, and support characteristics of the composite wing spar; the analysis and iterative design of the composite wing spar can be completed accurately and quickly; at the same time, the force transmission simulation of the composite wing spar of general aviation aircraft can be completed through the rough model, providing a reference for the simulation methods of other structures of general aviation aircraft (not limited to composite structures).

[0119] The above is only a preferred embodiment of the present invention. It should be pointed out that for ordinary technicians in this technical field, several improvements and modifications can be made without departing from the technical principles of the present invention. These changes and modifications should also be regarded as the scope of protection of the present invention.

Claims

1. A simulation method for composite wing spar structure of general aviation aircraft based on MSC.Nastran analysis and solution, characterized by: include: S1: Establish the finite element model of the entire wing; S2: Define the properties of the wing model structure, obtain the correct properties of the rod CROD unit, assign the correct properties to the finite element model, and establish the finite element model; When calculating the structural properties of the spar, each half of the spar cross section is divided into three regions: Region 1: the skin at the spar flange, Region 2: the skin boss and spar flange at the spar flange, and Region 3: the spar web. The skin boss and spar flange in Region 2 are constructed by layup and simulated using rod CROD elements. S3: Apply load case calculation to the established finite element model; extract the rod element force in region 2 from the calculation results, and use the correction factor to correct the rod CROD element load to obtain the actual load; The correction factor is h fem / h real , where h real h represents the distance between the neutral axis of the beam and the neutral axis of zone 2; fem represents the distance between the neutral axis of the beam and the position of the bar element; S4: Based on the cross-sectional stiffness ratio, the actual load is distributed to the skin boss and beam flange, and their respective strength checks are performed.

2. The method for simulating composite wing spar structures of general aviation aircraft based on MSC.Nastran analysis and solution according to claim 1 is characterized in that: The establishment of the overall finite element model of the wing as described in S1 includes: S11: extracting the outer surface of the skin from the digital model of the general aviation aircraft wing structure as a reference for constructing the finite element model, and adjusting the normal of the outer surface to point to the inside of the box segment; S12: Based on the benchmark, establish a finite element model of the composite wing spar and other structures of the general aviation aircraft; The Hypermesh software is used to establish a finite element model of the composite wing spar and other structures of a general aviation aircraft. The other structures described in S12 include wing spar strips, reinforcements, and wing spar webs. The wing spar strips and reinforcements are simulated using rod Crod units, and the wing spar webs are simulated using shell units Cquad4 and Ctria3.

3. The method for simulating composite wing spar structures of general aviation aircraft based on MSC.Nastran analysis and solution according to claim 1 is characterized in that: In addition to the wing spar structure, S2 also defines the properties of other wing model structures, including: the skin, ribs, and beam web structures are defined using the Pcomp paving method, and the layers are laid out according to the actual structural design situation.

4. The method for simulating composite wing spar structures of general aviation aircraft based on MSC.Nastran analysis and solution according to claim 1, characterized in that: In S2, the structural properties of the wing spar are calculated based on the principle of consistency of moment of inertia, including: S21: Calculate the distance between the neutral axis shared by the upper and lower regions 2 of the spar and the reference position as shown in formula (1): Where: A i — Areas of upper / lower skin bosses and upper / lower spar caps; E i ——Equivalent elastic modulus of the upper / lower skin bosses and upper / lower spar caps; c i — The distance between the upper / lower skin boss and the neutral axis of the upper / lower spar cap relative to the reference position; Among them, i=1 is the upper skin boss; i=2 is the upper edge of the beam; i=3 is the lower edge of the beam; i=4 is the lower skin boss; S22: The distance from the neutral axis of the upper and lower regions 2 of the spar to the reference position is calculated as follows: Where: A i — Areas of upper / lower skin bosses and upper / lower spar caps; E i ——Equivalent elastic modulus of the upper / lower skin bosses and upper / lower spar caps; c i — The distance between the upper / lower skin boss and the neutral axis of the upper / lower spar cap relative to the reference position; i=1 is the upper / lower skin boss; i=2 is the upper / lower spar cap; S23: Calculate the distance between the neutral axis of the upper and lower regions 2 and the shared neutral axis. The calculation formula is as follows: h real_upper =C real_upper -C g h real_lower =C g -C real_lower (3) Where: h real_upper ——The distance between the neutral axis of upper region 2 and the shared neutral axis; h real_lower - the distance between the neutral axis of lower region 2 and the shared neutral axis; C g - the distance between the neutral axis shared by the upper and lower parts of area 2 and the reference position; C real_upper - the distance from the neutral axis of the upper area 2 itself to the reference position; C real_lower - the distance from the neutral axis of the lower area 2 itself to the reference position; S24: According to equations (1)-(3), the area assigned to region 2 in the model is: Where: E fem ——The equivalent elastic modulus assigned to region 2 in the model; h fem —the distance between the neutral axis of the beam and the position of the rod element; h real — the distance between the neutral axis of the beam and the neutral axis of zone 2; E i_real ——True equivalent elastic modulus of the upper / lower skin bosses and upper / lower spar caps; A i — Areas of upper / lower skin bosses and upper / lower spar caps; S25: Modify the properties of the beam flange and skin boss. The modification principle includes: assuming that the moment of inertia of the finite element unit is equal to that of the actual structure, and simulating the action and load of the wing beam according to formula (5): Where: h real — the distance between the neutral axis of the beam and the neutral axis of zone 2; h fem —the distance between the neutral axis of the beam and the position of the rod element; A real ——The actual area of ​​Region 2; A fem —The area assigned to Region 2 in the model; E real ——The true equivalent elastic modulus of region 2; E fem ——The equivalent elastic modulus assigned to region 2 in the model.

5. The method for simulating composite wing spar structures of general aviation aircraft based on MSC.Nastran analysis and solution according to claim 1, characterized in that: In S3, MSC.Nastran software is used to apply load conditions to the finite element model.

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