Multi-perspective structural integrity-based computational investigations on airframe of Gyrodyne-configured multi-rotor UAV through coupled CFD and FEA approaches for various lightweight sandwich composites and alloys

1.1 Literature survey

The authors looked into the factors that contribute to deaths by drowning on beaches. A life-ring delivery drone system was proposed by the author to save time and lives. This gadget may potentially get a life ring to a person in trouble much quicker than lifeguards could. A stochastic simulation was run to calculate the victim’s location and the amount of time they had left to swim to safety before drowning. Based on the simulation results and the location of a nearby lifeguard tower, takeoff points for the drone were selected. Battery life is 20,000 mA·h, the octocopter weighs 10.2 kg, and the wheel base is 1,000 mm. This work includes conceptualization, mathematical analysis, simulations, a field study at Galveston Island’s beach, and a summary. So, the authors’ work also included a computational fluid dynamics (CFD) and structural analysis [1].

Their research assessed the viability of using UAVs in urban healthcare settings financially. It was suggested that emergency UAV transport to hospitals be modelled as a strategic undertaking to determine the best possible logistics. A unique facility placement algorithm was needed to determine what was needed to satisfy market demand. The author looked into the potential for using UAVs for intra-hospital delivery in National Health Service hospitals in London. Rather than focusing on improving range or payload capacity, the provided research emphasizes the need of strategically placing hubs across the network [2].

The authors have conducted extensive research on octocopters. They created a hybrid method that incorporates both CFD and experimental data with more traditional methods to improve the aerodynamic efficiency and performance of the drone in the low-speed subsonic zone. This research presented a new design to fix the problems with conventional monolayer and coaxial setups. The author built a model and analysed the aerodynamics of five octocopters in great detail. This was done using both the standard setup and the coaxial arrangement. It was found that the standard design had more drawbacks than the coaxial one. The author conducted a CFD analysis of rotor blades efficiency was crucial in achieving a brand-new configuration. The results showed that the novel configuration generated 71% more thrust than a traditional drone with the same fuselage arm length and 41.5% more thrust than a coaxially configured drone with the same blade length. The new configuration analysis proved also useful for other types of copters, such as hexacopters and octocopters [3].

The authors have conducted in-depth research on a quadcopter’s chassis. To carry out this research, the authors first had to meticulously build the quadcopter’s frame, then examine its static and dynamic characteristics, and then analyse its structure with a finite element (FE) code. The vertical thrust of the four motors was used to judge the structural integrity of the frame. The problems were fixed using E-glass fibre and polyamide nylon 6. The dynamic analysis was built on the premise of varying velocities and altitudes. The author recognized that results might differ depending on the motor and airframe used; however, they did remark that the major hub of the airframe was supported in a fixed position [4].

FE analysis (FEA) was applied to the wing and landing gear of the composite target drone. For this study, a 5 g symmetric pull-up force and a 1.5 g push-over stress were applied to the wing. The landing gear was analysed assuming a landing velocity of 1.4 m·s⁻¹ and a 2 g horizontal touchdown. The authors’ work used MSC/NASTRAN and LS-DYNA for this study. As part of a validation trial, the author subjected a prototype wing to a 6 g symmetric pull-up with sandbag loading. Invalidation results in 17% greater wing tip deflection than anticipated by FEA. The author attributes this variance to a defect in manufacturing. The sole substrate used in computer-driven experiments was E-glass. Moreover, a buckling analysis was carried out using the FE approach [5].

Development of the rotary-wing unmanned aircraft system (RUAS) was performed by the authors. The author examined the RUAS’s fuselage, main rotor, and tail rotor in great detail to get insight on methods for detecting and mitigating air pollution. All of the aerodynamic properties were calculated using CFD. Ansys Structural 17.2 was used for a lightweight composite material analysis, and Ansys preprocessing tool 17.2 was used for a composite material analysis, both of which were applied to the RUAS. Creating a surface model is the first step in the procedure’s workflow. After the framework is designed and built out of a variety of materials such as metal matrix composites and state-of-the-art hybrid composites, it is analysed to see how well it performs. Four different perspectives on the RUAV’s FSI (fluid–structure interaction) were considered: half fuselage, entire fuselage, main rotor, and overall. The results show that carbon-woven wet is the best material for building the RUAS under the given boundary condition [6]. Experiments verifying the model were carried out at the high-speed jet facility, grid convergence testing was carried out to guarantee a more precise result, and a comprehensive computation was supplied.

The authors have researched about the structural integrity and aerodynamic force variations of a vertical takeoff and landing (VTOL)-configured hybrid-blended wing body. Over the past 5 years, UAVs have been filling an increasingly important role in today’s society. UAVs can be customized for a wide variety of tracking, photography, and videography applications. The writers have a firm grasp on the characteristics of the perfect fixed-wing UAV (FWUAV). This project looked into the stability of FWUAV based on a blended wing body design from a computational aerodynamic, aero-structural, and control standpoint. For the Advanced Modelling Tool, the author focused on developing standard formulas [7].

The well-established computational coupled methodologies were used to investigate the disc brake of long-range UAVs. UAVs have seen explosive growth in popularity over the past decade. The proposed work posed a safety risk when assembling the necessary components for the procedure. The produced model was used in transient structural and thermal simulations, with the primary purpose of the study being to discover an appropriate material for UAV disc brakes. The author compared and contrasted the two seemingly identical strategies for estimating frictional force. Additionally, the author targeted their attention on the size, weight, and electronics of an FWUAV, which can carry a payload of 5 kg [8]. The effects of flexural behaviour on many high-tech aviation materials were investigated. The authors investigated cross-disciplinary enhancements. Selecting a material based on its intended application has led to the development of the ideal composite in recent years. Since vertical testing was necessary, an answer concerning deflection and slope was needed for the suggested task. The author spent much time and effort using a variety of cutting-edge techniques to examine the flexural properties of cutting-edge materials such as glass fibre-reinforced polymer (GFRP), Kevlar fibre-reinforced polymer (KFRP), and carbon fibre-reinforced polymer (CFRP). The primary focus of this work was on performing a thorough examination of the source material [9].

Using computational structural analysis, they looked into how to improve the frame of a multi-rotor UAV. Use of micro-electromechanical system components has allowed for a rapid technological advancement in UAVs in recent years. The +Frame, I-Frame, K-Frame, X-Frame, and Z-Frame were built as a consequence of computational structural analysis. Both boundary conditions on the “Z” frame and a grid convergence analysis on the “X” frame under maximum aerodynamic load were implemented. When tested under similar peak stress conditions, all three frame types (CFRP, GFRP, and KFRP) fared similarly well. The goal of this study was to determine the best quad airframe design for any given application and payload by comparing their respective structural and vibrational stability and strength [10].

The structural analysis of a rotor for an UAV was undertaken, and it was aimed to decrease the number of cracked propellers. Internal stress and subsequent propeller deformation were the focus of this work. The FSI analysis focused on the propeller contra designed in CATIA. The contra’s forward movement distance is 4.5 inches, and its diameter is 5 inches. The rotation’s effects were accurately estimated through to the discretizations in advanced sliding mesh. The coaxial propeller’s CFD simulation output was used to input the structural load. Analysis showed that using lightweight materials can reduce mass without compromising strength [11].

UAV wing materials, vehicle spoiler materials, and similar components can all benefit from the FE study conducted by the authors. In this case, CFRP is the material of choice, which is due to the fact that CFRP has more load-bearing capacity than the polymer matrix composite. As a reinforcing component, they use unidirectional carbon with a modulus of 230 GPa; however, they have also investigated the possibility of using materials with different orientations in the carbon layer. Epoxy resin serves as the glue in this situation. Forty computational specimens were chosen based on the flexural test and high life duration results, with CFRP-UD-230GPa-wet and CFRP-UD-230GPa-woven serving as the two primary specimens and the orientation angle of the fibres in the remaining 38 materials serving as the defining characteristic. A rectangular specimen was most frequently used because it required only a single-end support while being constructed. The rectangular specimen also had a 100 N vertical load applied to it. The best materials were produced when the orientation angle was 0 degrees for 8 layers and 90 degrees for 8 layers [12].

According to the authors, the outcomes of numerical simulations can be tightly regulated by actual boundary conditions, such as dependable load circumstances, ideal supports, and pure mechanical properties. The focus was on performing impact testing on carbon, glass, and Kevlar fibre composites using FSI analysis. Ansys Design Modeller was used to create the conceptual design, and the ACP software was used to create many composites. The optimal material for impact applications was determined when a battery of impact testing was performed to compare different materials. Given the evolving nature of the problem, a rigorous engineering approach was necessary for investigation [13].

According to studies conducted by the authors, UAVs have a wide variety of current applications, including agriculture, disaster management, military applications, and search and rescue. The propulsion system was vital to the operation of UAVs like the EADS Talarion that were meant to operate at medium altitudes for extended periods. Compact propulsion systems often make use of small gas turbine engines. Axial flow compressors are the standard in gas turbine engines. Blades can rotate or remain stationary during the stator and rotor stages, respectively. The aircraft industry is already seeing the benefits of composites’ promising future [14]. This is because composites have superior physical properties, such as higher specific strength, lower weight, and higher stiffness-to-density ratio

The authors used FSI analysis with a created FE model to perform the structural analysis of the wing to analyse the performance of composites in the skin. They used Selig 1223 model of the wing for the CFD analysis, and the pressure load obtained from the analysis was transferred to the created model of the composite wing. The authors estimated the design parameters of the wing to design the required computational model of the wing. Then, the wing underwent discretization to generate the FE model. The modelling of the composite wing section was created using ANSYS ACP and was integrated as the computational model for the CFD and FSI analysis. The FSI analysis was performed on composites with and without honeycomb structures. From the results of the analysis, they concluded that the composites with honeycomb provide better structural integrity to the wing skin on the basis of structural strength and stiffness [15].

Similarly, relevant references are studied and so the useful boundary conditions were obtained [16–41]. Most of the studies were only incorporated the lightweight composite materials and alloys in the airframe of UAVs [16–41]. Therefore, this work was finalized to impose sandwich composite materials in order to reduce the weight factor for UAVs. Because of this sandwich-based finalization, the weight factor can save up to 50% than the conventional lightweight materials. Apart from the weight reduction factor, the structural withstanding behaviour is also playing an important role. Hence, this work has been developed and comprised more than 50 sandwich material combinations. Since the finalized lightweight materials were slightly higher in counts, the flexible as well as advanced methodology only can solve this focused aim on this proposed UAV’s airframe. Henceforth, FSI-based computational investigation was decided to impose in this work, and the relevant computational conditions were found out [16–41].

Due to the focused application, the payload is heavier than the conventional multi-rotor UAVs. So, the design of this multi-rotor UAV has been developed with the aim of an innovative configuration and standard analytical calculations. Under methodology sections, firstly, the design calculations are mentioned; secondly, the problem formulations involved in computational fluid dynamics and structural integrity are incorporated; thirdly, the computational refinement studies are provided. Under results and discussions sections, two different manoeuvrings such as VTOL and forward speed operations are discussed. Both these phases contain four sub-sections, namely, results of base composite materials, results of base aerospace alloys, results of sandwich composites, and results of advanced hybrid-sandwich composites.

2.1 Design of Gyrodyne airframe

2.1.1 Estimation of payload weight

The components and payload weight are predicted to be 25,000 g based on a literature review, and these components are as follows: petrol can, diesel can, engine oil, brake drum, radiator, gear oil, and brake pads. The correlation between payload and total mass is as follows:

(1) W Pl = 0.512 W O .

Eq. (1) explains the relationship between weight of the payload and overall weight [6,7,8,10,22,31,32,33,34]:

W O = 25 0.512 = 48.828125 kg .

A rough diagram of the Gyrodyne UAV is shown in Figure 1.

Figure 1

Proposed rough diagram of Gyrodyne-configured drone.

2.2 Estimation of VTOL propeller’s diameter

To find the thrust required by single propeller, the following equation is used:

The thrust-to-weight ratio is fixed as 1.5 because of the additional gust loads that can be raised from the working environments [6,7,8,10,22,31,32,33,34]. The number of propeller short-listed for this configuration is 20. The requirement of thrust for single propeller in vertical direction is estimated as 3.66 kg. The standard formula for thrust is given in the following equation, which is taken from previously published studies [6,7,8,10,21,30,31,32,33]:

The aerodynamic property variations are not changed within 50 m tolerances in altitude of the picked working environment. So, the VTOL maximum velocity is picked as 25 m·s⁻¹ based on the mission requirement. The average velocity of the working fluid at the dangerous environments has been extracted from the literature survey [6,7,8,10,22,31,32,33,34]:

0.5 × ( ρ ) × A rotor × [ ( 25 ) 2 − ( 4.34 ) 2 ] = 35.93 ⇒ D Propeller VTOL = 0.36 m .

2.3 Estimation of forward propeller’s diameter

The thrust required by single propeller in the forward direction is estimated using Eq. (2). The thrust-to-weight ratio is fixed as 1.5 because of the additional gust loads that can be raised from the working environments [6,7,8,10,22,31,32,33,34]. The number of propellers short-listed for this configuration is 2. The requirement of thrust for single propeller in the forward direction is estimated as 6.1 kg. The standard formula for thrust is given as Eq. (3), which is taken from previously published studies [6–8,10,22,31–34]:

0.5 × ( ρ = 1.102 ) × A rotor × [ ( 25 ) 2 − ( 4.34 ) 2 ] = 59.841 ⇒ D Propeller Forward = 0.48 m .

2.4 Design of fuselage

To find the length of the fuselage, Eqs. (4) and (5) are used:

Distance between hub of the left ducted propeller and left extreme end of fuselage = 1 × diameter of the vertical propeller.

Distance between hub of the rigth ducted propeller and right extreme end of fuselage = 1 × diameter of the vertical propeller.

Thus,

2.5 Design of connecting arms

From trigonometric,

cos ( 45 ° ) = AB BC ⇒ cos ( 45 ° ) = 0.25 × Length of payload bay Length of connecting arm .

To find the length of connecting arm, the following equation is used:

Length of Connecting arm = 0.382 m.

To find the diameter of connecting arm, the following equation is used:

Diameter of connecting arm = Length of connecting arm 10 ⇒ 0.0382 m .

With the help of Fusion, Solid Works, and design calculation, the 3D model of the Gyrodyne UAV airframe is modelled. Figure 2 shows the isometric view of 3D model of Gyrodyne UAV airframe.

Figure 2

3D model of Gyrodyne UAV without arm – isometric view.

2.6 Discretization

This work is carried out with the assistance of grid convergence study, which will be discussed further on forthcoming sections. Figure 3 demonstrates the discretized model of the Gyrodyne UAV. Discretization is an important step that must be completed in a computational model before moving on to simulations. It will determine the outcome of this airframe model based on the elements and nodes. For CFD problems, due to the presence of the complicated curvature, unstructured tetrahedral elements are used for the development of discretization. For FEA problems, hybrid mesh that comprises tetrahedral and hexahedral-based discretization process is executed. Fine discretization produces more reliable results than ordinary discretization. In order to achieve more effective problem solving at edges, the sizing of the mesh properties has been modified to include proximity and curvature in the size function. For fine mesh, the smoothing parameter has also been modified to have a higher value.

Figure 3

Discretized view of Gyrodyne UAV – side view.

For the purpose of structural testing, the solid model of the Gyrodyne UAV is transformed into a computational model, which is then further improved to a moulded computational model with a thickness of 10 mm by the addition of 10 layers, each of which has a thickness of 1 mm. Figures 4 and 5 illustrate the surface model of the Gyrodyne UAV.

Figure 4

Isometric view of Gyrodyne UAV airframe – surface model.

Figure 5

Internal view of Gyrodyne UAV airframe – surface model.

2.7 Boundary conditions

As part of the FSI process, the pressure obtained in the fluent solver is imported into the computational model as external load and the ducted parts of the Gyrodyne UAV are arrested with fixed support. Incompressible flow-based turbulence characteristic-associated computational investigations are carried out in ANSYS Fluent. In which, the inlet speed of 25 m·s⁻¹ with k-epsilon enhanced wall treatment-based turbulence model are predominant boundary and initial conditions. Since the proposed UAV is medium in design, the 100,000-range-based Reynolds number has been imposed with coupled scheme-associated pressure and velocity coupling approaches. The effect of the imported pressure on the structural airframe of the Gyrodyne UAV is depicted in Figure 6. The effect of the fixed support and pressure on the surface model of the Gyrodyne UAV is depicted in Figure 7. The pressure acting on the fluid domain is depicted in Figure 8. This in-depth study examined four distinct classes of materials: base composites, base alloys, sandwich composites, and hybrid-sandwich composites. Seventy distinct materials are used in the initial computational research based on VTOL and forward speed manoeuvrings. The material properties are extracted from the literature survey [31] and the material library of the imposed computational tool. For contacts, all the structures are joined together through bonded connections. Since the model is developed in ANSYS Composite preprocessor, the bonded connection is default option. The one-way coupling-based FSI investigations are executed in ANSYS Workbench 17.2 with inbuilt facility of the system coupling approach.

Figure 6

Imported pressure acting on Gyrodyne UAV on its solid domain.

Figure 7

Typical view of fixed support and imported pressure acting on Gyrodyne UAV.

Figure 8

Typical view of pressure acting on Gyrodyne UAV in the fluid domain.

2.8 Solver descriptions and governing equations

Since the load is a known quantity from the CFD solution, the stiffness approach is the most appropriate one for the FSI investigation, and the load is further subdivided into steady and transient types depending on the method used to solve the model, in which case the constant gradually distributed load method is preferred. The FEA-based structural solver uses a governing equation to find the solution to the model, with the governing equation changing depending on the type of model being solved (here, the beam model), as shown in the following equation:

Stress and strain according to the imposed model for strain are two more potential results. This model uses a large variety of materials that are both isotropic and orthotropic, and the solver relies on displacement–strain matrices, which can be orthotropic or isotropic depending on the materials used. The relationship between displacement and strain is depicted by Eq. (9). Orthotropic displacement–strain matrices are represented by Eqs. (10)–(12):

where {ε} is the strain matrix, [B] is the strain–displacement matrix, and [u] is the displacement matrix.

To find the stress–strain relationship, the following matrix is used:

where {u} is the displacement matrix, [B] is the strain displacement matrix, and [D] is the stress–strain relationship matrix.

2.9 Grid-independent studies

This focused computational refinement study is to identify the ideal operating circumstances by carrying out a grid convergence analysis at a particular number of mesh elements. This condition included the maximum pressure-based comprehensive variation and its refinement. The level of saturation is figured out by first assessing the grid’s convergence in a variety of different conditions. The results of this grid convergence approach are presented in Table 1, which presents the general findings.

A dependable working condition of maximum pressure can be achieved during the meshing process by varying the size of the mesh components as they are added. That is saturated in case 6, which has mesh elements with a size of 3,848,430, and the maximum pressure obtained in that case is 380.06 Pa. Figure 9 describes how the pressure changes as a result of the different mesh components.

Figure 9

Maximum pressure from grid convergence test.

2.10 Experimental and simulation validation

2.10.1 Validation studies using experimental test and computational analyses

The precision of the enforced sophisticated methods of computation using ANSYS Workbench 17.2 needs to be verified by validating the collected computational outputs using experimental testing. In this research, the outcomes of experiments carried out at a facility that featured a high-velocity jet and the results of necessary computational approaches are compared and contrasted. Figure 10 depicts a facility with a jet that travels at a high rate of speed. The imposed test objects made of CFRP and aluminium alloy are shown in Figures 11 and 12, respectively.

Figure 10

Test setup for high-pressure load development.

Figure 11

Test specimen before structural failure – aluminium alloy.

Figure 12

Test specimen before structural failure – CFRP.

Although the pressure that is present inside the path of the jet is substantially higher, the tests are carried out in an atmosphere that is representative of that which might be encountered outside of the route of the jet. Figure 13 shows the different equipment that is used throughout the experiment, as well as a specimen that is intended to be evaluated.

Figure 13

Test setup along with test specimen.

The intense aerodynamic force brought on by the 40 bar supplied high pressure causes the test specimen constructed of aluminium alloy to fail. On the other hand, the test specimen constructed of CFRP retains its structural integrity throughout the whole experiment. The typical structural breakdown of an aluminium alloy test specimen is seen in Figure 14, which provides a representation of this breakdown. Based on the findings of the previous experiments, it is found that the woven CFRP had an ultimate stress of 3,500 MPa, whereas the improved aluminium alloy had an ultimate stress of 690 MPa. After successfully completing the experimental testing, the following step is to use ANSYS Workbench 17.2 to carry out the computations that are required for the tests to be considered valid. An accurate representation of experimental setup-based computational test specimen design data is generated using a recently developed computational platform of the high-speed jet route linked with the computational test specimen. This can be accomplished by establishing a connection between the high-speed jet route and the computational test specimen. The output of all of the completed computational models is shown in Figure 15, together with their respective input data. The structure of the computational model is depicted in a discretized form in Figure 16. Calculating the necessary pressure fluctuation on the test specimen is accomplished by the use of a computer simulation, which included pressure inlets and a discretized model. This is made possible through the simulation of the impact that the pressure would have on the test specimen. The typical distribution of the aerodynamic pressure both inside and above the test object is depicted in Figure 17. ANSYS Workbench 17.2 uses a sophisticated system coupling strategy in order to perform the process of transferring the aerodynamic pressure that is predicted on the model into the computationally produced models of aluminium alloy and CFRP-based composites. This task is necessary to complete the mission of transferring the pressure. The organized transfer of pressure that takes place is seen in Figure 18. When both models are exposed to imposed and transmitted aerodynamic loads, the FEA-based solver is used to compute the ultimate equivalent stresses for each model. Figures 19 and 20 exhibit the models created out of aluminium alloy and CFRP composite, respectively, in order to help demonstrate the structural consequences in a clearer manner.

Figure 14

Test specimen after structural failure.

Figure 15

A frontal view of a standard computational platform displaying the jet path of a high-speed test setup.

Figure 16

A zoomed 3-D elemental view of discretized structure.

Figure 17

A typical planar view-based representation of aerodynamic pressure variations in and over the computational test specimen.

Figure 18

Imported pressure on the computational test specimen.

Figure 19

Ultimate equivalent stress of aluminium alloy.

Figure 20

Ultimate equivalent stress of CFRP woven-based test specimen.

A maximum equivalent stress of 2853.3 MPa is predicted for the woven CFRP when this computational approach is used, whereas a maximum equivalent stress of 2157.7 MPa is calculated for the improved aluminium alloy. The yield stress on the aluminium alloy is determined with the assistance of the imposed advanced FSI method, and an acceptable level of stress is reached on the CFRP test specimen. In the experimental test setting, there are both a test specimen made of aluminium alloy that had fractured into pieces and a test specimen made of CFRP composite that is unbroken. Because of this, the decision has been made to impose specific computational approaches on the various components of the UAV, and one may rely on these techniques to deliver correct results. This is a result of the fact that the decision has been solidified into a choice.

2.10.2 Validation of FEM and computational analyses

As per an aforesaid boundary condition, the computational investigation on the shortlisted middle portion of the UAV’s airframe has been computed using FSI approach. Figure 21 shows the CFD pressure variations over the shortlisted section, and Figure 22 shows the outcome of the structural total deformation. Both Figures 21 and 22 are taken from the middle portion of the Gyrodyne UAV, which goes through the same FSI procedure as the rest of the model. The VTOL 15 m·s⁻¹ condition is being applied here, and the material being used is an aluminium alloy with ten layers, each of which is one millimetre in thickness.

Figure 21

Pressure contour on the mid-part of Gyrodyne UAV’s airframe.

Figure 22

Total deformation of the mid-part of Gyrodyne UAV’s airframe.

The mid-part of Gyrodyne UAV is solved as a beam with a midpoint load where the load is from Fluent solver, and the both ends are made to the fixed support. Figure 23 shows the 2D view of the solving perspective view of Gyrodyne UAV.

Figure 23

Typical boundary condition acting on the mid-part of Gyrodyne UAV.

The area of upper surface is 0.26517 m². As we know that, Force = Area × pressure ⇒ W = 0.26517 × 132.807 ⇒ 35.216 N. The FE equation (FEE) for the first element (AB) is expressed in Eq. (14), and also, the unique force relationship of the first element-based FEE is expressed in Eq. (15):

(14) F 1 M 1 F 2 M 2 = EI L 3 12 6 L − 12 6 L 6 L 4 L 2 − 6 L 2 L 2 − 12 − 6 L 12 − 6 L 6 L 2 L 2 − 6 L 4 L 2 y 1 θ 1 y 2 θ 2 ,

where E = 71 GPa and I = b d 3 12 ⇒ I = 0.99 × 10 − 6 m 4 ⇒ L = 0.3 m .

(15) F 1 M 1 F 2 M 2 = 0 − WL / 4 − WL / 2 0 ⇒ F 1 M 1 F 2 M 2 = 0 − 5.2824 − 10.5648 0 ,

0 − 5.2824 − 10.5648 0 = 2603.333 × 10 3 12 1.8 − 12 1.8 1.8 0.108 − 1.8 0.054 − 12 − 1.8 12 − 1.8 1.8 0.054 − 1.8 0.108 y 1 θ 1 y 2 θ 2 .

The FEE for second element (BC) is expressed in Eq. (16), and also, the unique force relationship of the second element-based FEE is expressed in Eq. (17):

(16) F 2 M 2 F 3 M 3 = EI L 3 12 6 L − 12 6 L 6 L 4 L 2 − 6 L 2 L 2 − 12 − 6 L 12 − 6 L 6 L 2 L 2 − 6 L 4 L 2 y 2 θ 2 y 3 θ 3 ,

(17) F 1 M 1 F 2 M 2 = − W / 2 0 0 − WL / 4 ⇒ F 1 M 1 F 2 M 2 = 17.608 0 0 − 5.2824 ,

17.608 0 0 − 5.2824 = 2603.333 × 10 3 12 1.8 − 12 1.8 1.8 0.108 − 1.8 0.054 − 12 − 1.8 12 − 1.8 1.8 0.054 − 1.8 0.108 y 2 θ 2 y 3 θ 3 .

After eliminating y 1 , θ 1 , y 3 , and θ 3 , because of the fixed support, the matrix form is 2 × 2 from 6 × 6 :

F 2 M 2 = 2603.333 × 10 3 24 0 0 0.216 y 2 θ 2 ⇒ 7.0432 0 = 2603.333 × 10 3 24 0 0 0.216 y 2 θ 2

7.0432 2603.333 × 10 3 = 24 y 2 ⇒ = 1.12727 × 10 − 7 m ⇒ y 2 = 1.12727 × 10 − 4 mm .

Using conventional FEE approach, the mid-span deflection of the short-listed as well as half-portioned UAV’s airframe is calculated. The comprehensive deflections under the approaches of computational and analytical methods are listed in Table 2. Additionally, the error percentage between these aforesaid methods is determined, which is within the acceptable limit. The obtained error rate is 3.01% only, which is very less value so the imposed computational procedure is, furthermore, verified and also proved that the computational procedure having the ability to provide reliable outcomes.

Table 2

Deflection error comparison

Deflection on body computational (mm)	0.00017871
Deflection on body analytical (cumulative) (mm)	0.0001092115
Difference in error (%)	3.11

3.1 FSI on Gyrodyne UAV – VTOL

The FSI investigations on Gyrodyne UAV are computed for two conditions based on their major manoeuvrings such as VTOL and forward conditions. For sample representations, Figures 24–28 show how deformation, stresses, and strain act on the airframe of Gyrodyne UAV under VTOL condition. With the aerodynamic load imported from the CFD analysis to this FSI analysis as pressure load, the materials imposed to the UAV’s airframe will begin to resist the load and deforms. The imposed pressure load acts through stresses such as normal and shear planes, which causes it to deform from its original shape creating strain on the airframe of the Gyrodyne UAV.

Figure 24

A typical structural outcome of CFRP-Wn-P-230GPa-based airframe – total deformation.

Figure 25

A typical structural outcome of CFRP-Wn-P-230-GPa-based airframe – equivalent stress.

Figure 26

A typical structural outcome of CFRP-Wn-P-230-GPa-based airframe – normal stress.

Figure 27

A typical structural outcome of CFRP-Wn-P-230-GPa-based airframe – shear stress.

Figure 28

A typical structural outcome of CFRP-Wn-P-230-GPa-based airframe – elastic strain.

Totally, four different material families, namely, base composites, base alloys, sandwich composites, and hybrid-sandwich composites, are investigated for various comprehensive investigations. Initially, the VTOL manoeuvring-based computational investigations are executed with 70 different materials counts.

3.1.1 Base materials – comprehensive report

Figures 29–38 show how total deformations, stresses, and strains are acting on the airframe of Gyrodyne UAV on various composite materials and alloys. The best material from all of these analysed cases is chosen based on the factors of lowest total deformation of the material, lowest normal and shear stresses acting on the material, and lower equivalent elastic strain values.

Figure 29

Comprehensive report of total deformation of various base composites.

Figure 30

Comprehensive report of equivalent stress of various base composites.

Figure 31

Comprehensive report of normal stress of various base composites.

Figure 32

Comprehensive report of shear stress of various base composites.

Figure 33

Comprehensive report of elastic strain of various base composites.

Figure 34

Comprehensive report of total deformation of various alloys.

Figure 35

Comprehensive report of equivalent stress of various alloys.

Figure 36

Comprehensive report of normal stress of various alloys.

Figure 37

Comprehensive report of shear stress of various alloys.

Figure 38

Comprehensive report of elastic strain of various alloys.

The comprehensive investigations are computed on base materials that undergone structural analysis for VTOL manoeuvring and so found out the best materials based on low reaction rate in terms of total deformation, equivalent stress, and other related properties. From low reaction rate, 8 materials are sorted out from 21 materials that are epoxy carbon UD (230 GPa) wet, epoxy carbon woven (230 GPa) Prepreg, Kevlar 49 UD, FR-4-based GFRP composite, grey cast iron, magnesium alloy, aluminium alloy, and copper alloy.

3.1.2 Sandwich materials – comprehensive results

The sandwich materials are formed from the short-listed base materials and incorporated with core materials such as honeycomb, polyvinyl chloride (PVC) foam, and styrene acrylonitrile (SAN) foam. Figures 39–53 show how total deformations, stresses, and strains are acting on the airframe of Gyrodyne UAV on various aforesaid sandwich materials.

Figure 39

Comprehensive report of total deformation of various sandwich materials.

Figure 40

Comprehensive report of equivalent stress of various sandwich materials.

Figure 41

Comprehensive report of normal stress of various sandwich materials.

Figure 42

Comprehensive report of shear stress of various sandwich materials.

Figure 43

Comprehensive report of equivalent elastic strain of various sandwich materials.

Figure 44

Comprehensive report of total deformation of various sandwich materials.

Figure 45

Comprehensive report of equivalent stress of various sandwich materials.

Figure 46

Comprehensive report of normal stress of various sandwich materials.

Figure 47

Comprehensive report of shear stress of various sandwich materials.

Figure 48

Comprehensive report of equivalent elastic strain of various sandwich materials.

Figure 49

Comprehensive report of total deformation of various sandwich materials.

Figure 50

Comprehensive report of equivalent stress of various sandwich materials.

Figure 51

Comprehensive report of normal stress of various sandwich materials.

Figure 52

Comprehensive report of shear stress of various sandwich materials.

Figure 53

Comprehensive report of equivalent elastic strain of various sandwich materials.

In a similar vein, a thorough examination of sandwich materials was conducted, focusing on a narrowed selection of eight base materials. Through this analysis, the optimal materials were determined based on their low reaction rate in relation to total deformation, equivalent stress, and other pertinent attributes. From the comprehensive results, the four best materials are sorted out: epoxy carbon woven (230 GPa) Prepreg, Kevlar 49 UD, aluminium alloy, and copper alloy.

3.1.3 Hybrid sandwiches

The hybrid-sandwich materials are formed using the best resultant materials from sandwich materials in various combinations that are incorporated with honeycomb, SAN foam, and PVC foam. Figures 54–68 show how total deformations, stresses, and strains are acting on the airframe of Gyrodyne UAV under VTOL manoeuvring condition on various composite materials and alloy-based sandwiches.

Figure 54

Comprehensive report of total deformation of various hybrid-sandwich materials.

Figure 55

Comprehensive report of equivalent stress of various hybrid-sandwich materials.

Figure 56

Comprehensive report of normal stress of various hybrid-sandwich materials.

Figure 57

Comprehensive report of shear stress of various hybrid-sandwich materials.

Figure 58

Comprehensive report of equivalent elastic strain of various hybrid-sandwich materials.

Figure 59

Comprehensive report of total deformation of various hybrid-sandwich materials.

Figure 60

Comprehensive report of equivalent stress of various hybrid-sandwich materials.

Figure 61

Comprehensive report of normal stress of various hybrid-sandwich materials.

Figure 62

Comprehensive report of shear stress of various hybrid-sandwich materials.

Figure 63

Comprehensive report of equivalent elastic strain of various hybrid-sandwich materials.

Figure 64

Comprehensive report of total deformation of various hybrid-sandwich materials.

Figure 65

Comprehensive report of equivalent stress of various hybrid-sandwich materials.

Figure 66

Comprehensive report of normal stress of various hybrid-sandwich materials.

Figure 67

Comprehensive report of shear stress of various hybrid-sandwich materials.

Figure 68

Comprehensive report of equivalent elastic strain of various hybrid-sandwich materials.

The best hybrid-sandwich materials are identified based on their low response rate in terms of total deformation, equivalent stress, and other features. This research included the four sandwich materials that had previously been narrowed down. Two materials are sorted out, which are (copper alloy)–honeycomb–(aluminium alloy) and (CFRP-woven-Prepreg-230-GPa)–honeycomb–(Kevlar 49 UD). These above-mentioned materials have lesser deformation and can withstand high stresses and strains caused by heavier payloads.

3.2 FSI on Gyrodyne UAV – forward

Comprehensive structural investigations are conducted on the next essential manoeuvre, forward speed operation, after the material withstanding investigations are completed successfully during VTOL manoeuvring. Deformation, stresses, and strains exerted on the airframe of a Gyrodyne UAV in response to a forward-speed-based boundary condition are depicted in Figures 69–73.

Figure 69

A typical structural outcome of KFRP-49 UD-based airframe – total deformation.

Figure 70

A typical structural outcome of KFRP-49 UD-based airframe – equivalent stress.

Figure 71

A typical structural outcome of KFRP-49 UD-based airframe – normal stress.

Figure 72

A typical structural outcome of KFRP-49 UD-based airframe – shear stress.

Figure 73

A typical structural outcome of KFRP-49 UD-based airframe – elastic strain.

3.2.1 Base materials

In line with the detailed investigations conducted on VTOL speed conditions, the forward speed operation is likewise evaluated using the four lightweight material families. Initially, a thorough examination was conducted on base composites, encompassing nine materials sourced from two distinct families, namely, CFRP and GFRP composites. Furthermore, a thorough analysis is conducted on base alloys, using a total of five different materials. The entire results of these two families of lightweight materials are depicted in Figures 74–82.

Figure 74

Comprehensive report of total deformation of various base composites.

Figure 75

Comprehensive report of total deformation of various alloys.

Figure 76

Comprehensive report of equivalent stress of various base composites.

Figure 77

Comprehensive report of equivalent stress of various alloys.

Figure 78

Comprehensive report of normal stress of various base composites.

Figure 79

Comprehensive report of normal stress of alloys.

Figure 80

Comprehensive report of shear stress of various base composites.

Figure 81

Comprehensive report of shear stress of various alloys.

Figure 82

Comprehensive report of equivalent elastic strain of various base composite materials.

The forward study involved a thorough examination of base materials, including composites and alloys. Subsequently, the optimal materials are identified based on their low response rate, as shown by measures such as total deformation, equivalent stress, and other relevant parameters. In a comparable manner, a total of eight lightweight materials have been selected from a pool of 21 materials. The materials that have made it to the final list include epoxy carbon UD 230 GPa wet, epoxy carbon woven 230 GPa Prepreg, Kevlar 49 UD, a woven composite based on FR-4 GFRP, grey cast iron, magnesium alloy, aluminium alloy, and copper alloy.

3.2.2 Sandwich materials – comprehensive results

The structural integrity of the sandwich materials is assessed using sophisticated computational systems. Figures 83–97 depict the comprehensive deformation, stresses, and strain exerted on the airframe of the Gyrodyne UAV. These measurements are obtained from several sandwich materials that are carefully selected and evaluated. The materials under consideration include honeycomb, PVC foam, and SAN foam.

Figure 83

Comprehensive report of total deformation of various sandwich materials.

Figure 84

Comprehensive report of equivalent stress of various sandwich materials.

Figure 85

Comprehensive report of normal stress of various sandwich materials.

Figure 86

Comprehensive report of shear stress of various sandwich materials.

Figure 87

Comprehensive report of equivalent elastic strain of various sandwich materials.

Figure 88

Comprehensive report of total deformation of various sandwich materials.

Figure 89

Comprehensive report of equivalent stress of various sandwich materials.

Figure 90

Comprehensive report of normal stress of various sandwich materials.

Figure 91

Comprehensive report of shear stress of various sandwich materials.

Figure 92

Comprehensive report of equivalent elastic strain of various sandwich materials.

Figure 93

Comprehensive report of total deformation of various sandwich materials.

Figure 94

Comprehensive report of equivalent stress of various sandwich materials.

Figure 95

Comprehensive report of normal stress of various sandwich materials.

Figure 96

Comprehensive report of shear stress of various sandwich materials.

Figure 97

Comprehensive report of equivalent elastic strain of various sandwich materials.

Similarly, the eight shortlisted base materials are used to conduct a thorough analysis on sandwich materials, uncovering the best materials based on a low response rate in terms of total deformation, equivalent stress, and other attributes. Four lightweight materials, namely, epoxy carbon woven (230 GPa) Prepreg, FR-4, magnesium alloy, and copper alloy, are settled on after careful consideration of the data.

3.2.3 Hybrid-sandwich materials

The sandwich materials with hybrid composition were fabricated and subjected to thorough examinations during the forward speed operation. The complete structural outcomes are depicted in Figures 98–112. Figures 98–112 depict the comprehensive analysis of total deformation, tension, and strain exerted on the airframe of the Gyrodyne UAV across several composite material sandwiches and alloy sandwiches.

Figure 98

Comprehensive report of total deformation of various hybrid-sandwich materials.

Figure 99

Comprehensive report of equivalent stress of various hybrid-sandwich materials.

Figure 100

Comprehensive report of normal stress of various hybrid-sandwich materials.

Figure 101

Comprehensive report of shear stress of various hybrid-sandwich materials.

Figure 102

Comprehensive report of equivalent elastic strain of various hybrid-sandwich materials.

Figure 103

Comprehensive report of total deformation of various hybrid-sandwich materials.

Figure 104

Comprehensive report of equivalent stress of various hybrid-sandwich materials.

Figure 105

Comprehensive report of normal stress of various hybrid-sandwich materials.

Figure 106

Comprehensive report of shear stress of various hybrid-sandwich materials.

Figure 107

Comprehensive report of equivalent elastic strain of various hybrid-sandwich materials.

Figure 108

Comprehensive report of total deformation of various hybrid-sandwich materials.

Figure 109

Comprehensive report of equivalent stress of various hybrid-sandwich materials.

Figure 110

Comprehensive report of normal stress of various hybrid-sandwich materials.

Figure 111

Comprehensive report of shear stress of various hybrid-sandwich materials.

Figure 112

Comprehensive report of equivalent elastic strain of various hybrid-sandwich materials.

In a similar manner, a detailed analysis on hybrid-sandwich materials is carried out based on the four sandwich materials that are nominated. As a result, the best materials are identified in terms of their low response rate in terms of total deformation, equivalent stress, and other associated attributes. After sorting through all four materials, the best two materials are shortlisted: (copper alloy)–honeycomb–(magnesium alloy) and (CFRP-woven-Prepreg-230-GPa)–honeycomb–(FR-4). These two materials make up the final product. These previously described materials have a lower potential for deformation and are able to bear the stresses and strains that are brought on by greater payloads.