The advancement in today's material science has driven composite materials to globally use in aircraft design with its superiority in high structural stiffness and significant weight reduction. Composite structures ordinarily comprise of laminates with various fiber orientation angles offers unique outcome, hence lead to optimized design for composite structure. The paper deals with the layerwise finite element model for static structural analysis of a CFRP laminated composite of unmanned aerial vehicle (UAV) wing. The objective of this study is to compare the results for different orientation of ply combinations which contributed to the high performance of composite materials that exhibit both orthotropic strength and stiffness properties. Both properties present unique challenges for analysis and design. The study is further up to determine the optimum design for selected ply combination on a wing with a tubercle design at the leading edge of the wing. Tubercles mimicking the protuberances on the leading edge of a Humpback whale pectoral flipper, offering great performance from an aerodynamic perspective. Hence, optimum design of composite is found from the tabulated stress and displacement for each ply combination, where the tubercles design at the leading edge of UAV wing showed better performance with a reduction in 38.75% of deformation and 46.83% of stress, compared to normal leading edge of NACA4415 airfoil.

Composite materials offer excellent strength-to-weight ratio with greater manufacturing feasibility of complex parts, unique contours and special feature, especially in the aircraft applications [1–3]. One of the basic components of an aircraft that received attention in the structural application of composite materials is the profile of the wings. Namely, one can predicts the behavior of physical systems under external influences of any product design's structure by using the finite element method of numerical solutions subjected to abstract equations of calculus [4]. The structural analysis based on the finite element method is known to be very effective numerical simulation and optimization method in the field of aerospace engineering. Previous studies related to finite element analysis presented in terms of software used and the outcome of the analyses conducted encompasses the scope of unmanned aerial vehicle (UAV) wing.

Sullivan et al. [5] used ABAQUS software to predict the potential failure of an ultralight UAV wing under critical in-flight loading conditions of pressure load. It was observed that the deflections occurred, whereby the wing failed at the limit load of 450kg with applied load of 136kg, 125kg, 106kg and 82kg on each wing at right and left side. However, significant difference of 1.2% between the experimental and computational data due to the thickness of the resin, which subjective to be included in the simulation. Kanesan et al. [6] also used ABAQUS to investigate the composite wing deflection for NACA4415 airfoil at different locations. The static analysis was carried out by applying distributed pressure load on the bottom skin that directly on top of main and aft spars at three different distances. Higher deflection is observed near the tip of the wing, which resulted within the range of 0.35–16.4% compared to bending experiment. Meanwhile, Shabeer and Murtaza [7] investigated the optimal design structure of UAV wing of all internal components and wing skin by using NASTRAN software. In this case, the wing skin was made of composite materials, whereas the other internal components were made of alloys. Similar analysis was carried out with different orientation of composite skin plies, whereby the sequence of [0/90/+45/−45/90/0] shown better performance in terms of displacement and maximum Von mises at the root of the wing. On the other hand, Kavya and Reddy [8] performed the static, modal and buckling analyses on three different composite materials including with/without the presence of spars and ribs by using ANSYS software. The study concluded that adding spars and ribs contributed to the high strength with minimum weight of the composite material of S Glass. Prabhu et al. [9] performed the static analysis on UAV wing of taper (NACA0012) and rectangular (NACA2412) using ANSYS software. In this case, the analyses were carried out by varying materials composition (alloy and composites) and different number internal components of the wing (stringers and ribs). Hence, the study revealed that stringers and ribs contributes to the stress action on the entire wing structure. Recent study by Chauhan et al. [10] performed the structural design process for UAV wing and subsequent optimization. The authors used mathematical modeling by MATLAB to calculate the loading as input, so that the static structural and buckling analysis can be performed using ANSYS. The spars were made of aluminum alloys, while ribs are made of balsawood. Equivalent von Mises stress was observed at the joint between spar and rib located near to wing root. Hence, the optimization in terms of weight was obtained by changing the position of spar, which at 22% from the chord, that contributed to the increase ratio for rib spacing and reduced total weight.

In practice, most special man-made composites consist of several elements of material known as ‘matrix’, reinforcement and core in form of fiber to increase the strength and stiffness of the structure [11–13]. Carbon fiber reinforced polymer (CFRP) is widely known as lightweight and strong materials that comprises of various type of fibers like carbon, glass and aramid as the reinforcement that are embedded in a polymer-based resin as the matrix [14,15]. The cloth formed (reinforcement and matrix) can be molded in a double curvature that required for aerodynamic shapes. Interestingly, core materials for sandwich structure is physically able to separate strength, and potentially in transmitting shearing forces across the sandwich [16,17]. Common types of cores used in aircraft applications are foam, honeycomb and wood. Therefore, the combination of reinforcement, matrix (resins) and core systems lead to the production of laminate that either can enhance or degrade the constituent material properties [18]. Hence, these combinations required to design a laminate of stacking plies of multiple layers composites to understand the structural behavior with unique orientations, as in Fig. 1.

Composite laminate structure [19].

The design goals of composite structure is its excellent of less in weight and high basic firmness in strength which characterized in terms of specific strength and specific modulus. Mechanically, the composite itself depend on the fiber properties and the degree of transmitted load on fiber–matrix bond. Thus, the deformation pattern in matrix surrounding a fiber is subjected to applied loads [18,20]. Besides that, the orientation of the fibers significantly influenced the strength of composite. In more specific, the properties of fibers direction such as continuous aligned, discontinuous aligned or random aligned may affect the stress–strain behaviors of fibers and matrix phases, the direction of load applied and the phase volume fraction. Thus, these circumstances may cause the composite failure like fracture. However, the inevitable challenges of optimizing the weight design of UAV composite with respect to its strength properties is the assignment of composite in terms of stacking sequence, fibers orientation and loading direction. In addition, this is due to the lack accurate data and the use of more familiar metals compared to composite [6,21]. In ANSYS, the composite fabrics orthotropic properties mainly to visualize the off axis stiffness to be expected from the fabric type [22]. The individual fabrics or even the combined stack ups of multiple plies that being applied to the structure in form of polar properties may reduce the trial and error at the analysis stage. The polar properties of fabrics may consist of three types namely unidirectional [0]n, bi-directional [0,90]n and quasi isotropic [0, ±45,90], as shown in Fig. 2.

Polar properties of fabric (a) unidirectional, (b) bidirectional, (c) quasi-isotropic [22].

In today's aeronautical applications, the advancement of flow control devices used on airfoils, wings and hydrofoils found in the natural world namely tubercles is one of the great factor contributing to increase stability, efficiency and less operational cost in the development of UAV industry. There has been a growing interest in researches of tubercles due to its prominent effects on the performance of aircraft wings such as more gradual stall, increase angle of stall, increase maximum lift, decrease the lift gradient near stall and many more [23–27]. Until recently, there are numbers of researches including numerical, experimental and analytical works that have been conducted to investigate the effects of TLE on the performance of humpback whale flippers [23,26,28–36]. Yet, most of these research works studied only on the aerodynamic performance of lift and drag and stall characteristics subjected to external flow.

Therefore, the aim of this paper is to investigate the structural performance of laminated composite at various fiber orientation angles and compare the advanced materials used in terms of optimal design of UAV wing. The wing design involves the design calculations of selected airfoil of NACA4415 including the external and internal members of wings, i.e. skin, spar, and ribs as well as the wing loading characteristics. The design is carried out using SolidWork and the assignment of composite as well as the analysis of structural deformation and applied loading conditions are done with the help of ANSYS 16.1 Workbench.

In particular, this paper is structured into sections and sub-sections. The parametric layout of the wing is elaborated in the next section. The following subsections include the geometrical configuration and material assignation. The modeling of layered composite structure is presented Section 3. The structural analysis on different fiber orientation will be done in Section 4, and followed by the result of analysis in Section 5. Conclusion will be provided in Section 6.

2Parametric layout of wing structureIn this study, the commencement step of static structural analysis is parametric layout of geometrical wing structure. In this step, the geometrical design is prepared and imported to ANSYS 16.1 Workbench to undergo the mesh generation of the imported model. The modeling process of layering the composite laminates is in the subsequent step. The analysis is carried out in the final step, whereby the condition of structural analysis and failure criteria are assigned to determine the structural response. The overall process is presented in Fig. 3.

2.1Geometrical configurationIn this study, the wing of NACA 4415 airfoil was designed using SolidWork, so that more accurate drawing of the airfoil shape can be obtained. The major structural components of UAV wing are three parts of skin, two spars, two outboard ribs, six leading ribs, five middle ribs and seven trailing ribs. In this study, the span length of the wing is set to 5.1257m and the chord length is 0.5886m. The design of UAV NACA 4415 is depicted in Fig. 4.

In this case, the spars are designed in such a way that combined the L-shape and U-shape. The spar to skin, rib to skin, and spar to rib are attached with adhesive joints, whereby the tied surfaces are assumed to be zero thickness subjected to the tied constraint.

2.2Material assignationAll major structural components of NACA 4415 UAV wing are made of composite fiber except for the bracket at the spar. The constructions of the composite components mostly used are carbon fiber fabric, unidirectional carbon fiber, Kevlar, honeycomb core with epoxy resin as matrix. The details of the wing components elastic and strength properties of material used are given in Tables 1 and 2.

Elastic properties materials (Sources: [6]).

Material | Density (kg/m3) | Elastic modulus, E11 (GPa) | Elastic modulus, E22 (GPa) | Poisson ratio, v12 | Shear modulus, G12 (GPa) |
---|---|---|---|---|---|

Carbon fiber fabric/epoxy | 1600 | 70 | 70 | 0.1 | 5 |

Carbon unitape/epoxy | 1600 | 140 | 10 | 0.3 | 5 |

Kevlar/epoxy | 1400 | 78.5 | 5.52 | 0.34 | 2.07 |

Honeycomb | 48 | 128.7 | 12.6 | 0.261 | 1.6 |

Strength properties materials (Sources: [6]).

Material | Tensile strength in fiber direction, Xt (MPa) | Tensile strength in fiber direction, Xc (MPa) | Tensile strength in transverse direction, Yt (MPa) | Tensile strength in transverse direction, Yc (MPa) | Shear strength, S (MPa) |
---|---|---|---|---|---|

Carbon fiber fabric/epoxy | 600 | 570 | 600 | 570 | 90 |

Carbon unitape/epoxy | 1500 | 1200 | 50 | 250 | 70 |

Kevlar/epoxy | 1380 | 276 | 29.6 | 137.9 | 43.4 |

Honeycomb | 2.344 | 4.07 | 2.344 | 4.07 | 6 |

Finite element mesh is the key role in determining the accuracy of validating since the nodes generated defined the output criteria of the analysis [6]. Prior to generating mesh, it is important to ensure that the mesh should accurately represent the geometry of the computational domain and loads. Kanesan et al. [6] also stated that the mesh should not contain elements with very large aspect ratio. The mesh also should adequately represent the large displacement or stress gradient in the solution. The shell elements (SHELL181) are used for the analysis of composite shells or plates [37]. In the case of laminated shell, the orientation of each lamina is defined from the given rotation angle relative to orientation for the entire shell section. Moreover, the properties of each lamina are defined by the linear elastic behavior for lamina under plane stress condition. The grid independence test is carried out from varying number of elements considering 10k, 18k, 20k, 40k, 65k, 80k, 100k, 120k, 140k, 165k, 180k, 200k, 250k, 300k, and 800k. Throughout the dependency test, the 250k element mesh provided an accurate solution with the hexahedral mesh. The maximum skewness of mesh is in good quality of 0.8. The mesh element of UAV wing is depicted in Figs. 5 and 6.

3Modeling layered of composite structureIn ANSYS [22], ACP (ANSYS Composite PrePost) module is used to apply composite materials of varying thickness and angles. This is applied only to the surface of wing skin, spars, and ribs. By using this module, the model developed in from of shell element of SHELL181 is assigned as single sheet of a specified thickness. The composite fabrics are created along with other composites and core materials, which are subsequently combined together to form a laminated composite materials. The composite fabric is created according to its thicknesses, and attached to each other to create a stack-up at assigned orientation (°). Then, to create a laminate, the stack-up with oriented 0° is sandwiched between the top and bottom plies. The most challenging is the complication arises in the design configuration with combination of various materials with different plies and various orientation in a 360° manner. This is important to adequately defined the model configuration with local coordinate systems (x1, y1, z1) at assigned orientation, including the other plate stacking of off-axis plies in relation to its appropriate local coordinate system (x3, y3, z3) [38]. This crucial consideration is well explained in Fig. 7.

Fiber orientation in three planar surface areas [38].

Referring to Fig. 3, the laminated composite is prepared according to the classical laminate theory (CLT) to investigate the behavior of composite sandwich structure subjected to external influence of load applied. In this case, the Kirchoff-Love hypothesis is applied in CLT approach, whereby the applied loads acting on the laminate structure is related to the mid-plane strains and curvature. In order to calculate the stress and strain in each lamina, both mid-plane and curvatures are assumed to be constant across the thickness. Yet, several assumptions are given [39,40]:

- •
Each lamina is orthotropic and quasi-homogeneous.

- •
Deformation are continuous and small through the laminate.

- •
A line, which is straight and perpendicular to the middle surface remains straight and perpendicular to the middle surface during deformation.

- •
The laminate is thin and it is loaded in its plane (plane stress), whereby out-of-plane (normal) direct stress is zero.

- •
Layers are perfectly bonded together and no slip occurs between the lamina interfaces.

- •
In-plane stress and curvature are small.

- •
Strain–displacement and stress–strain relationship are linear.

- •
Transverse shear strain (γxz andγyz) are negligible.

- •
Transverse normal strain ɛz is negligible compared to the in-plane strains ɛx and ɛy.

In ANSYS, the model of wing structure is interpreted in form of shell element (lamina). At first, the lamina is defined in terms of fabric, material and its thickness. As in Fig. 3, the subsequent step is the definition of rosette as reference direction of 0o of fiber direction, oriented selection set for layup direction and composite layup in modeling ply groups.

Theoretically, a laminate comprises number of plies, n and each ply has its thicknesses of tk. The thickness of the laminate, h can be derived as

The location of the mid-plane is h/2 from the top and the bottom surface of the laminate. The z-coordinates of each ply, k surface are given by:

The laminate created is depicted in Fig. 8.

Throughout the thickness of the laminate, integration of the global stresses in each individual ply provides cross-sectional forces and moments per unit length in x–y plane. Hence, cross-sectional forces and moments in relation to the mid-plane strains and curvature, is as follows [41–43]:

wherewhere Nx and Ny are normal force per unit length; Nxy is shear force per unit length. Mx and My are bending moment per unit length; Nxy is twisting moment per unit length.

Hence, the matrix of [A], [B] and [D] represent the respective extensional, coupling and bending stiffness of laminate. The [A] relates the resultant in-plane forces to the in-plane strains, while [D] relates the resultant bending moment to the curvature of laminate ply. [B] provides coupling effects of both forces and moments.

4Finite element structural analysis4.1Boundary conditionReferring to Mehta and Joshi [4], the known conditions such as force or displacement degrees of freedom at some nodal points were assigned in the finite element analysis as the boundary condition of the model. In this analysis, two important boundary conditions needed to be specified. The first is the attachment of the bracket to the fuselage. The bottom surface of the bracket was defined as fixed due to the fuselage not included in the model. All the displacements and rotations of the bottom surface of the brackets were set to zero. The second is the symmetrical condition of the wings that were defined as the boundary condition. The spars of the wing were set to be symmetrical about the YZ-plane. In this case, both conditions is defined to be fixed support points at UX=UY=UZ=ROTX=ROTY=ROTZ=0.

4.2Loading conditionIn this study, the calculation of the loads on the wing is subjected to the pressure that is distributed at the bottom skin of the wing, as studied by Kanesan et al. [6]. The directions inferred as the arrow labels in the nodal coordinate system, where FX=FY=0 and FZ=F. The concentrated load is specified directly on top of main and aft spars. The loading conditions are subjected to corresponding pressure values at four different locations, as shown in Table 3 and Fig. 9. In this case, the pressure was assigned based on the distance from X-axis, from the root of the wing.

Loading conditions for UAV composite wing.

Total load (kg) | Spar | Distance from the root of lower skin (cm) | |||||
---|---|---|---|---|---|---|---|

0–30 | 30–60 | 60–90 | 90–120 | 120–150 | 150–180 | ||

Pressure (Pa) | |||||||

40 | MS | 444.44 | 388.89 | 277.78 | 222.22 | 166.67 | 55.56 |

AS | 166.67 | 166.67 | 111.11 | 111.11 | 55.56 | 55.56 | |

41 | MS | 444.44 | 388.89 | 333.33 | 222.22 | 166.67 | 55.56 |

AS | 166.67 | 166.67 | 111.11 | 111.11 | 55.56 | 55.56 | |

42 | MS | 444.44 | 388.89 | 333.33 | 277.78 | 166.67 | 55.56 |

AS | 166.67 | 166.67 | 111.11 | 111.11 | 55.56 | 55.56 | |

43 | MS | 444.44 | 388.89 | 333.33 | 333.33 | 166.67 | 55.56 |

AS | 166.67 | 166.67 | 111.11 | 111.11 | 55.56 | 55.56 | |

44 | MS | 444.44 | 388.89 | 333.33 | 333.33 | 166.67 | 55.56 |

AS | 222.22 | 166.67 | 111.11 | 111.11 | 55.56 | 55.56 | |

45 | MS | 444.44 | 444.44 | 333.33 | 333.33 | 166.67 | 55.56 |

AS | 222.22 | 166.67 | 111.11 | 111.11 | 55.56 | 55.56 |

The static structural is conducted and validated the current finite element (FE) study with the available paper of simulation and experimental by Kanesan et al. [6]. Then, two parametric studies are conducted to investigate the performance analysis of the laminated composite wing, which is ply orientation and tubercles design at the leading edge.

4.3.1Parametric study I: laminate cases at different ply orientation for normal wingFirstly, the laminate cases of different ply orientation are carried out. In this case, five types of ply orientation are prepared in order to determine the significant of structural response under given physical loading conditions. The first study presented the variation of ply sequence in the stackup oriented subject to relative angle of 0°. Then, the sub-laminated is the sequence of plies based on fabrics and stacked at relative angles between the top and bottom plies. The ply orientation is tabulated in Table 4.

4.3.2Parametric study II: laminate cases at different ply orientation for spherical pattern at the leading edge of wing skinThe second is the different design of the wing skin, particularly at the leading edge of the wing. The wing is designed using the same airfoil with addition of spherical pattern subjected to amplitude and wavelength. Similar process as in Section 4.3.1, whereby the structural response of stress and deformation is determined to critically assess and compare the performance of ply orientation at different design. Since Refs. [35,44,45] proved better aerodynamic performance in CFD analysis, the optimum amplitude of A=0.025c and wavelength, λ=0.25c will be used in this study.

4.4Failure analysisIn ANSYS, the failure mode is subjected to strain and stress, material directions, principal directions or tension and compression. While the failure criteria is defined as maximum strain, maximum stress, Tsai-Wu, Tsai Hill, Hashin, Puck, etc. [46]. Those criteria can be classified in terms of safety based on the accordant weighing factor of each failure mode. In this study, Tsai-Wu failure criterion is applied to check the ultimate strength of a composite structure element that can be written as [47]:

where, σ1 is stress in fiber direction, σ2 is stress in transverse direction and τ12 is shear stress, C and T is the compression and tensile, respectively. Using the notion of first ply-failure, if the ply under consideration has failed, then Eq. (9) is violated. It is notable that both stresses can be distinguished subjected to appropriate coefficients. Noteworthy, it also can be easily incorporated in automated computational procedures.Furthermore, the strength ratio (S.R.) also can be evaluated in accordance with the Tsai-Wu failure theory (Eq. (13)) to examine the first lamina fails. The requirement verification of the structure is based on the value of Margin of Safety (M.o.S) subjected to the strength ratio. Fig. 10 depicts the process of determining the failure of structure.

5Results and discussion5.1Validation with simulation and experimental of bending test byPrior to further analyzing the performance of laminated composite, it is important to validate the current finite element (FE) analysis conducted with available paper from [6]. Table 5 shows the comparison of deflection results in simulation and experiment of bending test with current FE analysis.

Deflection results based on total load in simulation and experiment of bending test.

Total load (kg) | Deflection result (mm) | Difference (%) | Deflection result mm) | Difference (%) | |
---|---|---|---|---|---|

Simulation [6] | Current FE analysis | Experimental [6] | |||

40 | 12.666 | 12.095 | 4.51 | 13.530 | 10.61 |

41 | 12.930 | 12.311 | 4.79 | 13.860 | 11.18 |

42 | 13.330 | 12.628 | 5.27 | 14.140 | 10.69 |

43 | 13.727 | 12.945 | 4.22 | 14.840 | 12.77 |

44 | 13.799 | 13.019 | 5.65 | 15.220 | 14.46 |

45 | 13.953 | 13.148 | 5.77 | 15.820 | 16.89 |

From both results in Table 5, it is observed that the difference of deflection results is within the range of 4–17%. Autio et al. [48] stated that the error results given by commercial finite element method program should be less than 20%. Nurhaniza et al. [49] also supported that the range of 10–25% of comparison results from other finite element software. Hence, the comparison results proved that the current FE analysis developed is acceptable due to the value achieved for the displacement. Therefore, the obtained composite stack-ups of the FE analysis can be used for further optimization of the wing structure by changing the orientation of composite skin plies.

5.2Parametric study IAfter the validation, the highest load of 45kg is used to determine its performance by changing the ply orientations of the skin plies in ANSYS ACP module. The similar loading of 45kg as in Table 3 is applied on the bottom skin directly on top of main spar and aft spar. In this study, the analysis is conducted on five models of laminated composite wing skin, as in Table 4. Each ply sequence has superior polar properties that represent the stiffness and strength in the direction to the fibers. The polar properties of ply sequence are based on elastic modulus and shear modulus in accordance with the layup production plies, as shown in Table 6.

From the observation, the reinforcement directions of the lamina of plies, 0° and 90° present the maximum stiffness, whereas the maximum of 45° is the shear modulus. According to [50], the unidirectional of [0,0]n ply orientations may provide axial strength and stiffness, but it is more susceptible to damage. Meanwhile, the ply orientation of [0,90]n may provide transverse strength and stiffness, which is more damage tolerant than [0,0]n. The plies with only 0° and 90° may lead to delamination due to Poisson's mismatch. On the other hand, the ply orientation of ±45° provide shear strength and stiffness, with the combination with 0° or 90° which is more damage tolerant and more damage resistant. Thus, the load may carry similar oriented stackup within the laminate and help to avoid any delamination under load. Due to the anisotropic properties of lamina, these stackups are commonly overlapped along different directions in order to obtain a laminate with a quasi-isotropic behavior. Model 5 with ply sequence of stackup [0,0,0] and sub-laminate of [0,45,90] showed the significant structural performance as the lowest displacement and stress, as compared with other sequences, as depicts in Fig. 11(a) and (b). Furthermore, Model 5 also showed its margin of safety more than 0 and satisfies the ratio of material strength to design load by having the positive value of 0.875. According to Ref. [51], if the margin is higher than 0 with positive value indicated that the model can withstand additional load, which is twofold than the design load.

From the observation, the displacement of composite UAV NACA4415 occurred at the tip of the wing, whereas the stress is found nearer to the root of the wing. In this case, Model 5 of sublaminate [0,45,90] showed the lowest displacement and stress on the skin plies compared to the similar oriented sublaminated-plies of [0,0,0]. Though the ply sequence of [0,45,0] has small difference in displacement, the maximum shear stress of [0,45,90] showed the lowest result. The construction of the orientation composite plies showed the unique effect on structural performance of the wing, which is acceptable to withstand the design load given.

5.3Parametric study IIFrom the study of laminate cases at different ply orientation for normal wing, the study is further up on the orientation for spherical pattern at the leading edge of wing skin. The cloth formed from the laminated composite is molded in a double curvature that follow the spherical shapes in form of pattern along the leading edge of the wing. Since both designs applied the similar modeling of layered composite materials, the tubercles also shared the similar polar properties as in Fig. 12(a)–(e) subjected to respective models. The result of parametric study II is in Table 7.

From the result of different orientation of composite laminates on tubercles at the leading edge of wing, Model 5 showed the lowest value of total displacement with 7.903mm, with lower value of maximum shear stress compared to others. Similar as normal wing, the orientation of quasi-isotropic behavior proved that the combination of 0°, 45° and 90° is significantly affect the strength of the wing structure. Previous researches [7,52,53] also supported that quasi-isotropic layup proved its superiority in stiffness and strength in variation of direction of axial, transverse and shear. Hence, the obtained value of Model 5 is found to be within the limit of M.o.S of 0.875 in positive value, whereby the stresses at the wing skin indicate that the structural component of the wing is reliable and safe.

Fig. 13(a) and (b) depict both results of total displacement and maximum shear stress for tubercles wing, whereby the displacement is found at the wing tip and the stress occurred at the wing root. Most of the displacement and stresses for all models occurred at the same location, but differ in terms of its performance to withstand the design load. The orientation of [0°/(0°/45°/90°)/0°] showed better structural performances, as compared to other models.

Nonetheless, from the performance analyses of composite ply orientation between normal and tubercles wing, the Model 5 of [0°/(0°/45°/90°)/0°] of both designed is selected to determine the optimal design for UAV wing. The result is tabulated in Table 8.

Structural performance between normal and tubercles wing of NACA4415.

Design/configurations/structural parameters | Normal wing | Tubercles wing | Difference (%) |
---|---|---|---|

Total displacement (mm) | 12.903 | 7.903 | 38.75 |

Maximum stress (MPa) | 73.088 | 38.861 | 46.83 |

Shear elastic strain (mm/mm) | 0.0017243 | 0.002157 | 25.11 |

Margin of safety (M.o.S) | 0.875 | 0.125 | – |

From the observation, the maximum stress acting tangent to the skin of the normal wing is 73.088MPa. However, 45.83% reduction of maximum stress is obtained if the similar orientation applied on the tubercles design at the leading edge. On top of that, the tubercles design also showed a reduction of 38.75% difference of total displacement as compared to normal wing. Interestingly, tubercles design also showed an increase value of 25.11% difference of shear elastic strain. The margin of safety observed for both conditions under static loading are less than 1, where tubercles has positive value of 0.125 and normal wing stated of 0.875 with positive value. This is a reasonable safety as it proves that the wing will not fail under the load.

In overall, various factor contributed to the difference of structural performance between both design conditions. The variation in fiber orientation at the same skin thickness may lead to variation in the maximum stress, which is either increase or decrease. This is due to technique of stacking the ply, whereby the similarly oriented plies responsive to damage subjected to the load. The type of the material used in one laminate also plays important role in determining the structural response subjected to the loading. For instance, the laminate with more than two type of composite materials is more complex in determining its significant effect on the strength compared to the laminate with less than two materials. Therefore, the implication of composite failure criterion for strength performance is important to significantly different lamina strength whereby the structural geometry needs to acknowledge the margin of safety accordingly.

6ConclusionThe influence of the ply orientation on the composite laminate of NACA4415 UAV wing is evaluated. The analytic predictions show that the stress state that develops in variation fiber orientation with angle-ply has a significant effect on the strength. Two different design at the leading edge of the wing is presented, which is normal wing with clean airfoil at the leading edge and tubercle wing with spherical pattern at the leading edge. From the finite element analysis, the study is validated with the available research paper of the same airfoil of NACA4415 comparing both simulation and experiment. The current study of finite element analysis showed the obtained value of total displacement is in acceptable range. Hence, the selected loading condition is chose to further the study toward determining the performance analysis under variation of ply orientation. The obtained structural performance from the normal wing is compared with the tubercle wing. Both conditions proved that the plies orientation of [0°/(0°/45°/90°)/0°] from Model 5 has better structural performance compared to other. Meanwhile, from the optimal design perspective, wing with tubercle design at the leading edge showcased a reduction of 38.75% and 45.83% difference of total displacement and maximum stress, respectively. For further study, the wing with tubercles design at the leading edge, whereby the integration of manufacturing cost into the wing structural optimization can be improved from the determination of optimal set of minimum structural weight and manufacturing cost by incorporating several optimization procedures such as the parametric geometry definition, generating 3D CAD model, generating finite element method, structural dimension optimization, cost estimation and layout optimization, with the aid of computational software.

Conflict of interestThe author declares no conflicts of interest.

This work is supported by UPM under GP-IPS grant, 9647200. The authors would like to express their gratitude and sincere appreciation to Department of Aerospace Engineering, Faculty of Engineering, Universiti Putra Malaysia and Laboratory of Biocomposite Technology, Institute of Tropical Forestry and Forest Products (INTROP), Universiti Putra Malaysia (HICOE) for the close collaboration in this work.