Parametric Optimization of Load Bearing Elements for Composite Wing of an Unmanned Aerial Vehicle

This paper describes the optimizing results of structural elements of the composite wing of an unmanned aerial vehicle. The thickness and composite lay-up structure of load-bearing elements and wing skin are determined using the ANSYS software package. The optimal structure is presented using the Pareto set method of the "ideal center" basing on four criteria: minimum mass, deflection, normal stress, and maximum safety factor of the wing. Verification calculations were carried out to determine the safety factor of the load-bearing wing structure using a geometrically nonlinear model in FEMAP software.


Introduction
An unmanned aerial vehicle (UAV) is defined as a type of aircraft with an engine that does not carry a human pilot, uses aerodynamic force to create a lift, can fly autonomously or remotely, and can carry a payload. In recent decades, the use of UAV has been rapidly expanding in both civilian and military aviation. The weight of the vehicle is always an important parameter, characterizing the technical perfection of the UAV, as this characteristic depends on the maximum altitude and range of flight, as well as the handling of the aircraft and its maneuverability. Currently, most aircraft designs are made of layered composites due to their high specific strength and rigidity characteristics, as well as the relatively high level of technology for composites. This study looked at the two-spars straight wing structure with a wingspan of 10.5 m and a chord length of 0.78 m. The position of the front and rear spars was fixed at a relative distance of 30% and 60% of the chord from the edge of the wing. The distance between the ribs was 400 mm [1]. The purpose of this work was to determine the optimal design and structural parameters of the composite wing of an unmanned aerial vehicle.

Methodology
The main structural elements of composite wing were considered as the upper and lower skin, two I-beam spars and 15 ribs. Aerodynamic pressure load was distributed on the surface of the lower skin of the UAV wing. As an optimization of the design parameters of the wing were considered in two stages. In the first stage, the thicknesses of the load-bearing elements and wing skin were determined using laminated composite materials: unidirectional CFRP, CFRP fabric and fiberglass. In the second stage, the optimization of the laying up structure of the optimal composite material was considered: CFRP fibric, considering the existing aerodynamic loads. The geometric model of the UAV wing is presented in Fig. 1. The basic technical requirements for the composite UAV wing were accepted in according to the standard of USAR 337. The estimated aerodynamic pressure was determined using the following equations [2,3]: Where q y -aerodynamic load acting on the wing; q w -structural load of the wing; q sresulting aerodynamic load of the wing; q ∑ -aerodynamic pressure acting on the wing; G 0 = 450 kg -UAV take-off mass; G w = 72 kg -wing mass, g -acceleration of free fall, () bzwing chord length n a max = 3.8 -coefficient of load factor; f = 1.5 -coefficient of safety factor.

Material properties
All the structural elements of UAV wing are made of composite materials layered structure. Three types of materials forming composite layers were considered: unidirectional carbon plastics, woven carbon plastics, and unidirectional E-glass plastic. The mechanical characteristics of these materials are given in the table. 1 [4].

Modeling and simulation calculation
The optimal thickness of the load-bearing elements was calculated using the OPTISLANG module in the ANSYS software package. The initial values of the thickness of the loadbearing elements and the wing skin were given in the same way for all materials according to Table 2. When choosing the optimal thickness, the wing deflection limit was taken at 5% of wingspans [5]. In the first step of optimization of the wing design elements, 400 variants for each material were considered. In the figures 2a, 2b and 2c on the plane of the criteria mass and deflection presents Pareto-optimal variants for each material, which are indicated: i g -variant No. i of fiberglass, j w -variant No. j of CFRP fabric, k u -variant No k of unidirectional CFRP. For each material, the optimal variant was chosen using the following Pareto set "ideal center" equation. In the next step, the remaining three variants 6 g , 54 w and 28 u were determined by the value of the third criteria, margin of safety in the action of aerodynamic pressure. Margin of safety calculations were made using a geometrically non-linear model in the FEMAP software package. Regarding the criteria for minimum mass and maximum margin of safety with using the additional criteria of "ideal center" the best variant was CFRP fabric 54 w (Fig. 3).  In table 3 shows the thickness of the structural elements of the composite wing for the optimal variant No. 54 w .

Optimization of composite lay-up structure for UAV wing
The layering angles were optimized using the OPTISLANG module in the ANSYS software complex. The angle of the layers for each constructive wing element ranged from 90 to -90 degrees, with an initial value of 0 degrees [6]. Since at this stage of optimization of the thickness of all the design elements have already been defined and thus the mass of the wing has been fixed, the minimum deflection and minimum normal stress criteria were adopted as optimization criteria. In the process of optimization, 210 variants of the composite lay-up structure were considered. The resulting Pareto-optimal variants are presented in Fig. 4 as points on the criteria plane. Results of verification calculations by margin of safety of these Pareto-optimal variants showed that only one of them, namely, variant No. 202 meets the specified requirements for the reserve of stability. This variant was chosen as the optimal one for all the criteria considered. The optimal angle of layers for all load-bearing elements and wing skin was obtained as 0, ±90 degrees.