# Design, Computational Aerodynamic, Aerostructural, and Control Stability Investigations of VTOL-Configured Hybrid Blended Wing Body-Based Unmanned Aerial Vehicle for Intruder Inspections

Authors: Vijayanandh Raja; Ramesh Murugesan; Senthil Kumar Solaiappan; Beena Stanislaus Arputharaj; Parvathy Rajendran; Hussein A. Z. AL-bonsrulah; Deepak Thakur; Abdul Razak; Abdulrajak Buradi; Abiot Ketema
Journal: International Journal of Aerospace Engineering
 (2023)

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## Abstract

Unmanned aerial vehicles (UAVs) are gaining in popularity and sophistication in today’s modern world. UAVs are now available in a wide range of configurations. A UAV’s many applications include aerial photography and videography and target tracking. The upward-pointing propellers of some modern fixed-wing UAVs make it possible for them to take off and land vertically. Surveillance and intruder inspections are two areas where the blended wing body (BWB) configuration shines. This is because its weight is spread uniformly throughout the body, its radar signal is weaker than that of alternative configurations, and there is a relatively small amount of interference with its movement. With common design factors in mind, like vertical takeoff and landing, aerodynamic drag, and fundamental wing stability, the optimal BWB plan form for surveillance is designed. CATIA is used to finish the conceptual design of the BWB-based UAV. A fluid-structure interaction (FSI) study is carried out after the model has been examined in ANSYS Fluent. The UAV’s responsiveness is improved through simulation in the MATLAB environment after a proportional-integral-derivative-type altitude controller was developed. The results demonstrate that providing the UAV with an altitude instruction enhances its performance. Given the flexibility of the suggested BWB UAV’s design, we have decided to limit its maximum forward speed to 75 m/s and its maximum rate of vertical ascension to 50 m/s. Rapid BWB UAVs like the one seen here are quite helpful in dangerous situations.


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## Body

## 1. Introduction

A fixed-wing unmanned aerial vehicle (FWUAV) might look like a miniaturized aeroplane. It has propellers that can be used to move forward or to take off and land vertically. FWUAVs can be controlled from the ground or by the computers on board if they have the right software. FWUAVs can fly because of how their wings are set up and how fast they move forward. Unmanned aerial vehicles (UAVs) with fixed wings that can move on their own have propellers that are powered by internal combustion engines or brushless DC motors. Gliders, on the other hand, are launched with a winch or pulled behind an aeroplane. With the ailerons, elevator, and rudder, the drone’s controller can change the roll, pitch, and yaw of the unmanned aircraft. Also, the rudder can be used to move the plane in any direction. Fixed-wing UAVs can get their power from batteries, gas or diesel engines, or other sources. Most UAVs have a fixed wing, a canard wing, a flying wing, or a body with a mix of these types of wings. There are also hybrid wing bodies and wings that can fly.The traditional UAV has wings that do not move and looks like a small aeroplane. It is moved by the tractor or pusher propellers. The vehicle needs a standard tail for pitching and yawing, a propeller for thrust, a rectangular wing for lift, and a fuselage body big enough to hold the payload. The vehicle gets lift from its rectangular wing. The canard wing UAV structure looks like a small version of a canard plane. Like a canard plane, it has a small wing in front of the main wing of a FWUAV. This layout is shown by canard wings. The UAV’s canard wing, on the other hand, saves space. They are taken into account because of how they affect flying control, instability, trim, and the flow over the main wing. These unmanned aerial vehicles are not used very often because simpler ones are easier to make and work better for the job. There are simpler designs because the cons are more important than the pros (UAVs). Flying wings can go farther and stay in the air longer than other unmanned aerial vehicles because their wings are bigger, their drag surfaces are smaller, and their batteries last longer. Unmanned aerial vehicles, on the other hand, have a traditional fuselage and tail surfaces that are flat. This unmanned aerial vehicle can be put together quickly by using spars to build the airfoil section and skin to cover it. The fuselage and the wing are both part of the same structure in a blended wing body design (BWB). It has a body in the middle, a wing, and a place where the two meet. The whole thing is made up of these three parts. At this point, the wing shape and the main body shape come together. This design has a low ratio of wetted area to internal volume, the possibility of an elliptical lift distribution, a cross section that changes smoothly, and room for an engine. Also, the cross section changes in a smooth way.


### 1.1. Literature Survey

This article compared the production ease, continuity, simplicity of arm liftoff, mass, and ground clearance of the three classic kinds of UAV architectures. Furthermore, the authors go over the pluses and minuses of each UAV design. Also, the authors focused on the ease of production, the consistency of the task, the simplicity of the arm liftoff, the weight, and the height above the ground. The authors additionally go over the advantages and disadvantages of each design, starting with the former. According to the literature [1], UAVs with a high wing are the most reliable option for use in life-or-death situations. Medium-range long-endurance unmanned aerial vehicle (MALE UAV) concept design with blended wing-body architecture is the focus of this article. In this regard, the airfoils FX 76-MP-120, MH-18, and MH-108 were chosen for the design, and CAD (computer-aided draft) software was used to create three different versions of the concept. After doing a CFD (computational fluid dynamics) performance analysis and a stability analysis, the team concluded that the wingtip tailplane configuration was the most effective overall [2]. Researchers arrived at this verdict after performing both sorts of analysis. The primary focus of this article is the procedure for enhancing the efficiency of UAVs with fixed-wing configurations. It also details the results of a wind tunnel experiment with a blended wing body UAV. The main goal of the test was to assess the level of stability. Results showed that, except from takeoff and landing, the UAV exhibited good stability throughout the flight [3]. Within the scope of this research, three distinct configurations for a blended wing body passenger aeroplane are examined and analyzed. The research article evaluates and ranks the three designs based on criteria like flight and field performance, static stability, balance, weight, and passenger accommodation, but it does not decide which one is the best. Some examples of such characteristics are as follows: some of these features are the ability to fly and perform on the field, as well as static stability, balance, weight, and passenger capacity. Also, consider static stability, balance, and weight [4]. This article takes a close look at the design and operation of a blended wing body aeroplane. The purpose of this research is to provide the findings of an aerodynamic analysis conducted on four potential blended wing body platform configurations. The primary focus of this study is on the effectiveness of the platforms for use in applications requiring unmanned aerial vehicles travelling at low subsonic speeds. The blended wing body extended from the centerline of the fuselage to the wingtip, and it was made up of three different airfoils; the CFD programme is used to analyze these and other design features of all four models. These specific airfoils were given the model names MH-91, E374, and NLF1015. Integrating the Reynolds-averaged-Navier-Stroke equations with a low Reynolds turbulence model yielded accurate shear and normal stress acting on the wing surfaces. The purpose of this action was to improve precision. This was done to provide the highest degree of precision in the final results [5–10]. The study [5] used a variety of turbulence models, including k-omega SST and Spalart-Allmaras, to analyze the external fluid’s flow and identify its features. To learn about the characteristics of the ambient fluid, this was done. This article explores the theoretical framework for, and the aerodynamic properties of, a VTOL-integrated blended wing body UAV. The essay is focused on the vertical takeoff and landing capabilities of the VTOL-integrated blended wing body UAV. This article focuses on the vertical takeoff and landing capabilities of the VTOL-integrated blended wing body UAV. This arrangement allows for vertical takeoff and landing by utilizing a propeller positioned in a duct along the plane’s midline. Located smack dab in the middle of the plane is this duct. It can take off and land vertically due to the design of the plane. Experiments were conducted on two versions of the same model to determine the thrust of four propellers hidden within the fuselage at various angles of attack. The same model was used for all of these tests. The device’s range and operational endurance are constrained by the capacity of the battery currently being used. [6]. According to the cited literature review and other online sources, one must conceptually construct three to four models by incorporating improvements from prior models, and one model must be optimized for surveillance application. This realization resulted from the realization that having a single model that is optimized for surveillance application is essential. This suggests the importance of having access to at least one model designed for use in such a setting as surveillance. This is the conclusion that may be derived from the need for a model optimized for surveillance applications. Inference reached from the data. When the first UAV plans were discovered, they were used to create the prototype for the BWB. With the use of relevant literature, researchers looked at the past connections between factors including payload weight, total weight, aspect ratio, and wingspan [2–5, 7, 11]. Roskam’s work [12], detailed in [6, 8, 9], served as the basis for the theoretical calculations. Since the majority of the existing literature relied on CFD simulations, we started with them as part of our testing strategy. Because it would yield the most precise findings, it was thought that the trial-and-error approach would be the ideal place to start with the design. It is expected that the procedure would go through three or four iterations before an appropriate design is found for use in surveillance applications. Those methods were used in the studies that are described in [3–6] and [7, 8].


## 1.1. Literature Survey

This article compared the production ease, continuity, simplicity of arm liftoff, mass, and ground clearance of the three classic kinds of UAV architectures. Furthermore, the authors go over the pluses and minuses of each UAV design. Also, the authors focused on the ease of production, the consistency of the task, the simplicity of the arm liftoff, the weight, and the height above the ground. The authors additionally go over the advantages and disadvantages of each design, starting with the former. According to the literature [1], UAVs with a high wing are the most reliable option for use in life-or-death situations. Medium-range long-endurance unmanned aerial vehicle (MALE UAV) concept design with blended wing-body architecture is the focus of this article. In this regard, the airfoils FX 76-MP-120, MH-18, and MH-108 were chosen for the design, and CAD (computer-aided draft) software was used to create three different versions of the concept. After doing a CFD (computational fluid dynamics) performance analysis and a stability analysis, the team concluded that the wingtip tailplane configuration was the most effective overall [2]. Researchers arrived at this verdict after performing both sorts of analysis. The primary focus of this article is the procedure for enhancing the efficiency of UAVs with fixed-wing configurations. It also details the results of a wind tunnel experiment with a blended wing body UAV. The main goal of the test was to assess the level of stability. Results showed that, except from takeoff and landing, the UAV exhibited good stability throughout the flight [3]. Within the scope of this research, three distinct configurations for a blended wing body passenger aeroplane are examined and analyzed. The research article evaluates and ranks the three designs based on criteria like flight and field performance, static stability, balance, weight, and passenger accommodation, but it does not decide which one is the best. Some examples of such characteristics are as follows: some of these features are the ability to fly and perform on the field, as well as static stability, balance, weight, and passenger capacity. Also, consider static stability, balance, and weight [4]. This article takes a close look at the design and operation of a blended wing body aeroplane. The purpose of this research is to provide the findings of an aerodynamic analysis conducted on four potential blended wing body platform configurations. The primary focus of this study is on the effectiveness of the platforms for use in applications requiring unmanned aerial vehicles travelling at low subsonic speeds. The blended wing body extended from the centerline of the fuselage to the wingtip, and it was made up of three different airfoils; the CFD programme is used to analyze these and other design features of all four models. These specific airfoils were given the model names MH-91, E374, and NLF1015. Integrating the Reynolds-averaged-Navier-Stroke equations with a low Reynolds turbulence model yielded accurate shear and normal stress acting on the wing surfaces. The purpose of this action was to improve precision. This was done to provide the highest degree of precision in the final results [5–10]. The study [5] used a variety of turbulence models, including k-omega SST and Spalart-Allmaras, to analyze the external fluid’s flow and identify its features. To learn about the characteristics of the ambient fluid, this was done. This article explores the theoretical framework for, and the aerodynamic properties of, a VTOL-integrated blended wing body UAV. The essay is focused on the vertical takeoff and landing capabilities of the VTOL-integrated blended wing body UAV. This article focuses on the vertical takeoff and landing capabilities of the VTOL-integrated blended wing body UAV. This arrangement allows for vertical takeoff and landing by utilizing a propeller positioned in a duct along the plane’s midline. Located smack dab in the middle of the plane is this duct. It can take off and land vertically due to the design of the plane. Experiments were conducted on two versions of the same model to determine the thrust of four propellers hidden within the fuselage at various angles of attack. The same model was used for all of these tests. The device’s range and operational endurance are constrained by the capacity of the battery currently being used. [6]. According to the cited literature review and other online sources, one must conceptually construct three to four models by incorporating improvements from prior models, and one model must be optimized for surveillance application. This realization resulted from the realization that having a single model that is optimized for surveillance application is essential. This suggests the importance of having access to at least one model designed for use in such a setting as surveillance. This is the conclusion that may be derived from the need for a model optimized for surveillance applications. Inference reached from the data. When the first UAV plans were discovered, they were used to create the prototype for the BWB. With the use of relevant literature, researchers looked at the past connections between factors including payload weight, total weight, aspect ratio, and wingspan [2–5, 7, 11]. Roskam’s work [12], detailed in [6, 8, 9], served as the basis for the theoretical calculations. Since the majority of the existing literature relied on CFD simulations, we started with them as part of our testing strategy. Because it would yield the most precise findings, it was thought that the trial-and-error approach would be the ideal place to start with the design. It is expected that the procedure would go through three or four iterations before an appropriate design is found for use in surveillance applications. Those methods were used in the studies that are described in [3–6] and [7, 8].


## 2. Proposed Methodology

The BWB-UAV is constructed with the assistance of a standardized formula as well as a sophisticated modeling tool. Sophisticated numerical modeling is utilized in order to simulate the actual working environment of the unmanned aerial vehicle in order to ascertain whether or not the BWB-modeled UAV is able to accomplish the desired missions while keeping expenses to a minimum.


### 2.1. Conceptual Design

The recommended methodology for this piece of research is fully dependent on the computational approach; as a result, it is important that the essential component, the BWB UAV, be modeled. An industry-standard formula is applied in order to determine the design parameters, components, boundary conditions, and specifications for a variety of BWB UAVs. Because of the intricate nature of the BWB UAV’s architecture, modeling plays a significant role in the process of developing the UAV’s conceptual design.


#### 2.1.1. Estimation of Overall Weight

The payload is a stabilized four-sensor UAV gimbal that can rotate a full 360 degrees, has a moving target indicator built in, and can record in 720p HD. The study assumes a payload weight of 5 kg, whereas the actual weight, including all accessories, is 4.5 kg. The total mass of the UAV can be roughly estimated by correlating past data with the weight of the payload. Thus, field work was done to refine payload weight estimates for the missions that made the cut. The new and distinctive link between payload weight and total weight of BWB UAV is derived from the established relationship provided by Equation (1). The typical comprehensive representations of link between payload mass and overall UAV mass are shown in Figure 1. According to the results of Equation (1), the expected payload mass is 5 kg and the predicted overall UAV mass is 16.5 kg.
(1)WPl=0.298096989⇒WO=50.298096989⇒WO=16.7731kg.Figure 1
Historical relationship between overall weight and payload weight.


#### 2.1.2. Design of Unique Propeller for Forward Speed

To design a high-performance propeller for forward motion, relevant researchers consider variables such overall weight, forward speed, propeller pitch, and propeller diameter. The fastest rate of progress that can be achieved in this endeavor is envisioned to be 75 m/s. The important relationships such as thrust and pitch of FWUAV’s forward speed propeller are expressed in Equations (2) and (4), respectively.
(2)T=0.5×ρ×A×VForwardspeed2−Vo2=100%oftheUAV′sweight2,(3)AreaForwardspeed=0.023831495m2⇒DiameterForwardspeed=6.86≈7inches,(4)Pitch=ForwardspeedofUAV’spropellerin“inch/second”Rotationalspeedofthepropellerin“RPS”,(5)Pitch=2952.76354.07=8.339≈8.35inches,(6)CL=2×Lρ×VForwardSpeed2×A.Lift, forward speed, flow density, and cross-sectional area of the propeller are the primary variables in Equation (6), which yields an estimated coefficient of lift of 0.997155556. Standard formulas, such as the one presented in Equation (7), are then used to determine the Reynolds number once an estimate has been made.
(7)Re=Densityofair×UAV’svelocity×diameterofthepropellerDynamicviscosityofair.The determined lift coefficient (CL) and the Reynolds number are both utilized in the process of calculating and comparing the drag coefficient (CD), for a variety of different NACA aerofoils. For the conceptual design of the propeller, the NACA 25112 aerofoil profile is utilized. This shape functioned with a lower coefficient of drag than the majority of other aerofoil profiles.
(8)Pitchangleθ=arctangentP2×π×r,(9)Chordlengthb=8×π×sinθ×tanθ−1/1.2×tanθ/1+1/1.2×tanθ×rn×CL.The chord length, pitch, and its placements in the propeller were computed for the forward speed’s propeller based on the aforementioned Equations (8) and (9). The values for these parameters are listed in Table 1. The forward speed executing propeller is illustrated in Figures 2 and 3, with constructions produced in accordance with the data that was determined.Table 1
Dimensions of forward speed propeller.
Location (mm)
Pitch angle (theta) (degree)
Chord length (mm)
8.71
75.51
16.22
17.42
62.67
24.08
26.13
52.22
26.97
34.84
44.06
27.26
43.55
37.75
26.35
52.26
32.83
24.96
60.97
28.94
23.46
69.68
25.82
21.98
78.39
23.27
20.59
87.10
21.16
19.32Figure 2
A typical design of unique propeller for forward speed maneuvering.Figure 3
Unique propeller for forward speed maneuvering.


#### 2.1.3. Design of Unique Propeller for VTOL Maneuvering

The key design parameters are estimated with the use of mathematical equations. The diameter, pitch, chord lengths, and pitch angles of the propeller, which is aimed at carrying out VTOL operation, are the details that are discussed. The atmospheric velocity is measured to be 4 m/s, and it is assumed that the maximum rate of ascension is 50 m/s.(10)0.5×ρ×AVTOL×VVTOLspeedofUAV2−Vo2=100%oftheBWBUAV’sweight2,(11)AreaPropellerforVTOL=0.151211375m2⇒DiameterPropellerforVTOL=17.275inches,(12)PitchPropellerforVTOL=0.6×diameterPropellerforVTOL,(13)PitchPropellerforVTOL=0.6∗17.275=10.365inches,(14)CL=2×LMax.atVTOLρ×VVTOLspeedofUAV2×AVTOL.The aforementioned equation (14) yields a lift coefficient of 0.982222222. After establishing this estimate, the Reynolds number is obtained by applying the common formulas provided in Equation (7). The VTOL operation propeller lacks the NACA 25112 aerofoil, which increases the forward propeller’s coefficient of drag over the VTOL operation propeller. The chord length, pitch, and locations in the propeller’s chords for forward speed were determined using Equations (8) and (9) previously mentioned. The important relationships such as thrust and pitch of FWUAV’s VTOL propeller are expressed in Equations (10) and (12), respectively. Table 2 contains the findings of this calculation. These authors are using a coaxial propeller technology, which enables them to design propellers that revolve both clockwise and anticlockwise using CAD design software. Figures 4 and 5 display, respectively, the CAD models of the VTOL propellers.Table 2
Dimensions of VTOL propeller.
Location (mm)
Pitch angle (theta) (degree)
Chord length (mm)
21.94
62.364
30.542
43.88
43.68
34.36
65.82
32.48
31.35
87.76
25.523
27.54
109.695
20.91
24.183
131.634
17.66
21.42
153.573
15.261
19.166
175.512
13.43
17.306
197.45
11.981
15.76
219.39
10.81
14.46Figure 4
Conceptual design of unique propeller for VTOL maneuvering.Figure 5
Unique propeller for VTOL maneuvering.


#### 2.1.4. Design of Primary Data of BWB UAV

The main design criteria that went into making the BWB UAV were its wingspan, length, planform area of the blended wing, and root chord. For these parameters to be found, a standard analytical method was used. Because of how important the wing loading is to the design process, we have decided that the wing loading for this BWB unmanned aerial vehicle should be 20 kg/m2.
(15)Swing=WOverallweightofBWBUAVW/SWingLoadingofBWBUAV,(16)Swing=16.773120=0.838655kg/m2.Equation (15) is used to calculate the planform area of this blended wing, and then, using precedents, the aspect ratio (AR) is assumed to be 4. The assumed AR has been imposed in Equation (17) that contains the link between wingspan and planform area of BWB UAV.
(17)AR=bwing2Swing,(18)Aspectratio=bwing20.838655⇒bwing=0.838655×4=1.832m.The typical relationship between length and wingspan of BWB-UAV is expressed in Equation (19). In most cases, the length-to-wingspan ratio of an unmanned aircraft falls around the range of 70–90%. The authors made the assumption that the design ratio (ƞ) for this work is 70%; therefore, if that is the case,
(19)L=ƞ×bwing,(20)LUAV=0.7×1.832=1.2824m.The relationship that exists between the chord length, wing area, and wing span is quite evident in Equation (21), and it can be observed rather simply.
(21)Swing=b×Cwing−root,(22)Cwing−root=0.8386551.832=0.458m.As a consequence of the analysis of the relevant literature, it was established that a taper ratio of 0.5 is the optimal number to employ in order to offer the lowest drag while maximizing lift at an angle of attack that is advantageous. As a direct consequence of this, the value of taper ratio that is chosen for this research is 0.5. The typical representation of taper ratio of BWB-UAV is expressed in the following equation:(23)Taperratioλ=CtCr.A high-wing design is envisioned for the purposes of this work. In order to compute chord at any span-wise position, the wingspan is always an essential aspect to take into consideration. With the assistance of these analytic approaches and a variety of additional supporting equations, the one-of-a-kind and highly advanced HBWB unmanned aerial vehicle can be constructed; the names of its components can be found in Figure6.Figure 6
Design of advanced HBWB UAV-typical isometric view.


### 2.2. Computational Fluid Dynamics (CFD)

The major objective of the cutting-edge and flexible CFD tool is to provide an accurate prediction of the aerodynamic pressures and forces (lift and drag) that the HBWB UAV will be subjected to in the foreseeable future. Within the domain of computational fluid dynamics, the discretization method is regarded as being of the utmost significance. In order to convert continuous functions into discrete functions, it is necessary to ensure that the distance between each point in the function remains unchanged during the transformation. If one makes the assumption that the flow’s maximum speeds are 75 m/s in the forward direction and 50 m/s when it is operating in VTOL mode, then one might be able to reach the conclusion that the flow does not compress. This immediately leads to the implementation of a solution that is based on the application of pressure. It is possible to use ANSYS Fluent to model the unmanned aerial vehicle that has been proposed in order to calculate the amount of drag it would generate while carrying out a range of maneuvers. Since the elaborate design leads to a rather high rate of turbulence formation, a traditionalk-epsilon turbulence model with quality of the second order is adopted. The working pressure is determined to be 101325 Pa in light of the facts presented here. Due to the fact that the maximum velocities were determined to be 75 and 50 m/s, respectively, the identical sets of inputs were applied for both sets of calculations. On the surface of the vehicle, there is not a problem with sliding, but free slip is present on the unmanned aerial vehicle. The wall is what governs the volume, and it has a set of dimensions that are predetermined and dictated by the geometry of a cylinder. The wingspan is ten times greater than the length of the fuselage, while the length of the fuselage is fifteen times greater than the length of the wings. Figure 7 not only gives a concise example of the border conditional components but also indicates that the control volume has the shape of a cylinder.Figure 7
Control volume of CFD problem with clear names.The SIMPLE approach is applied in order to derive the velocity and pressure coupling, and it is needed that all significant equations have a quality level of second order or higher. The flow is incapable of being compressed and keeps its density constant throughout its entirety, ultimately reaching a maximum finalized velocity of 75 m/s. Because this control volume has been determined, the pressure-based solution can now be used to create a forecast regarding the properties of the fluid. This prediction can be made because this control volume has been established. The primary fluid parameters, such as pressure and volume, are measured at each node of a control volume, and then, the values of those parameters are determined. These parameters include pressure and volume. When an unmanned aerial vehicle is equipped with a propeller, the risk that it may experience turbulence is significantly increased. As a direct consequence of this, we are going to make use of the turbulence that is present. The large amount of labour that would be required to adequately display the turbulent flow led to the selection of the second order as the best option to pursue.The development of a discretized unstructured grid on the BWB UAV is depicted in an accurate manner in Figure8. The identification of curvature and area reduction-based regions using BWB UAVs is primarily focused with the establishment of the grid. The process of precisely creating the grid is concluded when these facilities are utilized, and as a result, there is a large improvement in the possibility of achieving a trustworthy output.Figure 8
Discretized structure of the entire control volume–VTOL operation.


### 2.3. Fluid-Structure Interaction (FSI)

This imposed FSI puts a variety of materials through tests for forward speed and VTOL operations so that it can select the material that is most likely to be able to withstand the working environmental conditions that have been described. In order to manufacture UAVs, manufacturers need to use components that are primarily made of lightweight materials that also have a high load resistance. The incorporation of a variety of lightweight materials into structural calculations that are constrained to aerodynamic force imposing conditions allows this to be achieved. To make an estimate of the structural results, computations with the help of finite computational element analyses (FEA) are carried out. In this simulation, an unstructured mesh is required to be used. The fine mesh facility is utilized for the production of difficult components such as vertical and horizontal propellers. The entirety of the structure is compelled to make use of the mesh facilities, which are dependent on the curvature and proximity of the base object, i.e., UAV. After much deliberation, the final rating for the mesh quality was 0.984 out of 1. The intricate mesh of the BWB UAV is displayed in its entire splendor in Figure9.Figure 9
Discretized structure of a normal HBWB UAV.The most essential boundary conditions for this FSI analysis are the impact on the aerodynamic load, the fixed support, and the remote displacements, while the most important initial conditions for this FSI analysis are the mechanical properties of lightweight materials. As can be seen in Figure10, the aerodynamic load that is placed on the BWB UAV is completely taken into account and transferred using a one-way coupling FSI method.Figure 10
Load importation on advanced HBWB-VTOL maneuvering.The CFD method is utilized most frequently in the process of estimating and extracting the aerodynamic load that is placed on a BWB UAV. Aside from this UDL-based aerodynamic load, the remote displacement on the bother the propellers is excellently used to present the representation of rotary action of these proposed propellers, and support deformations are clearly involved in the representation of hold constrictions at the landing gear of BWB UAV.


## 2.1. Conceptual Design

The recommended methodology for this piece of research is fully dependent on the computational approach; as a result, it is important that the essential component, the BWB UAV, be modeled. An industry-standard formula is applied in order to determine the design parameters, components, boundary conditions, and specifications for a variety of BWB UAVs. Because of the intricate nature of the BWB UAV’s architecture, modeling plays a significant role in the process of developing the UAV’s conceptual design.


### 2.1.1. Estimation of Overall Weight

The payload is a stabilized four-sensor UAV gimbal that can rotate a full 360 degrees, has a moving target indicator built in, and can record in 720p HD. The study assumes a payload weight of 5 kg, whereas the actual weight, including all accessories, is 4.5 kg. The total mass of the UAV can be roughly estimated by correlating past data with the weight of the payload. Thus, field work was done to refine payload weight estimates for the missions that made the cut. The new and distinctive link between payload weight and total weight of BWB UAV is derived from the established relationship provided by Equation (1). The typical comprehensive representations of link between payload mass and overall UAV mass are shown in Figure 1. According to the results of Equation (1), the expected payload mass is 5 kg and the predicted overall UAV mass is 16.5 kg.
(1)WPl=0.298096989⇒WO=50.298096989⇒WO=16.7731kg.Figure 1
Historical relationship between overall weight and payload weight.


### 2.1.2. Design of Unique Propeller for Forward Speed

To design a high-performance propeller for forward motion, relevant researchers consider variables such overall weight, forward speed, propeller pitch, and propeller diameter. The fastest rate of progress that can be achieved in this endeavor is envisioned to be 75 m/s. The important relationships such as thrust and pitch of FWUAV’s forward speed propeller are expressed in Equations (2) and (4), respectively.
(2)T=0.5×ρ×A×VForwardspeed2−Vo2=100%oftheUAV′sweight2,(3)AreaForwardspeed=0.023831495m2⇒DiameterForwardspeed=6.86≈7inches,(4)Pitch=ForwardspeedofUAV’spropellerin“inch/second”Rotationalspeedofthepropellerin“RPS”,(5)Pitch=2952.76354.07=8.339≈8.35inches,(6)CL=2×Lρ×VForwardSpeed2×A.Lift, forward speed, flow density, and cross-sectional area of the propeller are the primary variables in Equation (6), which yields an estimated coefficient of lift of 0.997155556. Standard formulas, such as the one presented in Equation (7), are then used to determine the Reynolds number once an estimate has been made.
(7)Re=Densityofair×UAV’svelocity×diameterofthepropellerDynamicviscosityofair.The determined lift coefficient (CL) and the Reynolds number are both utilized in the process of calculating and comparing the drag coefficient (CD), for a variety of different NACA aerofoils. For the conceptual design of the propeller, the NACA 25112 aerofoil profile is utilized. This shape functioned with a lower coefficient of drag than the majority of other aerofoil profiles.
(8)Pitchangleθ=arctangentP2×π×r,(9)Chordlengthb=8×π×sinθ×tanθ−1/1.2×tanθ/1+1/1.2×tanθ×rn×CL.The chord length, pitch, and its placements in the propeller were computed for the forward speed’s propeller based on the aforementioned Equations (8) and (9). The values for these parameters are listed in Table 1. The forward speed executing propeller is illustrated in Figures 2 and 3, with constructions produced in accordance with the data that was determined.Table 1
Dimensions of forward speed propeller.
Location (mm)
Pitch angle (theta) (degree)
Chord length (mm)
8.71
75.51
16.22
17.42
62.67
24.08
26.13
52.22
26.97
34.84
44.06
27.26
43.55
37.75
26.35
52.26
32.83
24.96
60.97
28.94
23.46
69.68
25.82
21.98
78.39
23.27
20.59
87.10
21.16
19.32Figure 2
A typical design of unique propeller for forward speed maneuvering.Figure 3
Unique propeller for forward speed maneuvering.


### 2.1.3. Design of Unique Propeller for VTOL Maneuvering

The key design parameters are estimated with the use of mathematical equations. The diameter, pitch, chord lengths, and pitch angles of the propeller, which is aimed at carrying out VTOL operation, are the details that are discussed. The atmospheric velocity is measured to be 4 m/s, and it is assumed that the maximum rate of ascension is 50 m/s.(10)0.5×ρ×AVTOL×VVTOLspeedofUAV2−Vo2=100%oftheBWBUAV’sweight2,(11)AreaPropellerforVTOL=0.151211375m2⇒DiameterPropellerforVTOL=17.275inches,(12)PitchPropellerforVTOL=0.6×diameterPropellerforVTOL,(13)PitchPropellerforVTOL=0.6∗17.275=10.365inches,(14)CL=2×LMax.atVTOLρ×VVTOLspeedofUAV2×AVTOL.The aforementioned equation (14) yields a lift coefficient of 0.982222222. After establishing this estimate, the Reynolds number is obtained by applying the common formulas provided in Equation (7). The VTOL operation propeller lacks the NACA 25112 aerofoil, which increases the forward propeller’s coefficient of drag over the VTOL operation propeller. The chord length, pitch, and locations in the propeller’s chords for forward speed were determined using Equations (8) and (9) previously mentioned. The important relationships such as thrust and pitch of FWUAV’s VTOL propeller are expressed in Equations (10) and (12), respectively. Table 2 contains the findings of this calculation. These authors are using a coaxial propeller technology, which enables them to design propellers that revolve both clockwise and anticlockwise using CAD design software. Figures 4 and 5 display, respectively, the CAD models of the VTOL propellers.Table 2
Dimensions of VTOL propeller.
Location (mm)
Pitch angle (theta) (degree)
Chord length (mm)
21.94
62.364
30.542
43.88
43.68
34.36
65.82
32.48
31.35
87.76
25.523
27.54
109.695
20.91
24.183
131.634
17.66
21.42
153.573
15.261
19.166
175.512
13.43
17.306
197.45
11.981
15.76
219.39
10.81
14.46Figure 4
Conceptual design of unique propeller for VTOL maneuvering.Figure 5
Unique propeller for VTOL maneuvering.


### 2.1.4. Design of Primary Data of BWB UAV

The main design criteria that went into making the BWB UAV were its wingspan, length, planform area of the blended wing, and root chord. For these parameters to be found, a standard analytical method was used. Because of how important the wing loading is to the design process, we have decided that the wing loading for this BWB unmanned aerial vehicle should be 20 kg/m2.
(15)Swing=WOverallweightofBWBUAVW/SWingLoadingofBWBUAV,(16)Swing=16.773120=0.838655kg/m2.Equation (15) is used to calculate the planform area of this blended wing, and then, using precedents, the aspect ratio (AR) is assumed to be 4. The assumed AR has been imposed in Equation (17) that contains the link between wingspan and planform area of BWB UAV.
(17)AR=bwing2Swing,(18)Aspectratio=bwing20.838655⇒bwing=0.838655×4=1.832m.The typical relationship between length and wingspan of BWB-UAV is expressed in Equation (19). In most cases, the length-to-wingspan ratio of an unmanned aircraft falls around the range of 70–90%. The authors made the assumption that the design ratio (ƞ) for this work is 70%; therefore, if that is the case,
(19)L=ƞ×bwing,(20)LUAV=0.7×1.832=1.2824m.The relationship that exists between the chord length, wing area, and wing span is quite evident in Equation (21), and it can be observed rather simply.
(21)Swing=b×Cwing−root,(22)Cwing−root=0.8386551.832=0.458m.As a consequence of the analysis of the relevant literature, it was established that a taper ratio of 0.5 is the optimal number to employ in order to offer the lowest drag while maximizing lift at an angle of attack that is advantageous. As a direct consequence of this, the value of taper ratio that is chosen for this research is 0.5. The typical representation of taper ratio of BWB-UAV is expressed in the following equation:(23)Taperratioλ=CtCr.A high-wing design is envisioned for the purposes of this work. In order to compute chord at any span-wise position, the wingspan is always an essential aspect to take into consideration. With the assistance of these analytic approaches and a variety of additional supporting equations, the one-of-a-kind and highly advanced HBWB unmanned aerial vehicle can be constructed; the names of its components can be found in Figure6.Figure 6
Design of advanced HBWB UAV-typical isometric view.


## 2.1.1. Estimation of Overall Weight

The payload is a stabilized four-sensor UAV gimbal that can rotate a full 360 degrees, has a moving target indicator built in, and can record in 720p HD. The study assumes a payload weight of 5 kg, whereas the actual weight, including all accessories, is 4.5 kg. The total mass of the UAV can be roughly estimated by correlating past data with the weight of the payload. Thus, field work was done to refine payload weight estimates for the missions that made the cut. The new and distinctive link between payload weight and total weight of BWB UAV is derived from the established relationship provided by Equation (1). The typical comprehensive representations of link between payload mass and overall UAV mass are shown in Figure 1. According to the results of Equation (1), the expected payload mass is 5 kg and the predicted overall UAV mass is 16.5 kg.
(1)WPl=0.298096989⇒WO=50.298096989⇒WO=16.7731kg.Figure 1
Historical relationship between overall weight and payload weight.


## 2.1.2. Design of Unique Propeller for Forward Speed

To design a high-performance propeller for forward motion, relevant researchers consider variables such overall weight, forward speed, propeller pitch, and propeller diameter. The fastest rate of progress that can be achieved in this endeavor is envisioned to be 75 m/s. The important relationships such as thrust and pitch of FWUAV’s forward speed propeller are expressed in Equations (2) and (4), respectively.
(2)T=0.5×ρ×A×VForwardspeed2−Vo2=100%oftheUAV′sweight2,(3)AreaForwardspeed=0.023831495m2⇒DiameterForwardspeed=6.86≈7inches,(4)Pitch=ForwardspeedofUAV’spropellerin“inch/second”Rotationalspeedofthepropellerin“RPS”,(5)Pitch=2952.76354.07=8.339≈8.35inches,(6)CL=2×Lρ×VForwardSpeed2×A.Lift, forward speed, flow density, and cross-sectional area of the propeller are the primary variables in Equation (6), which yields an estimated coefficient of lift of 0.997155556. Standard formulas, such as the one presented in Equation (7), are then used to determine the Reynolds number once an estimate has been made.
(7)Re=Densityofair×UAV’svelocity×diameterofthepropellerDynamicviscosityofair.The determined lift coefficient (CL) and the Reynolds number are both utilized in the process of calculating and comparing the drag coefficient (CD), for a variety of different NACA aerofoils. For the conceptual design of the propeller, the NACA 25112 aerofoil profile is utilized. This shape functioned with a lower coefficient of drag than the majority of other aerofoil profiles.
(8)Pitchangleθ=arctangentP2×π×r,(9)Chordlengthb=8×π×sinθ×tanθ−1/1.2×tanθ/1+1/1.2×tanθ×rn×CL.The chord length, pitch, and its placements in the propeller were computed for the forward speed’s propeller based on the aforementioned Equations (8) and (9). The values for these parameters are listed in Table 1. The forward speed executing propeller is illustrated in Figures 2 and 3, with constructions produced in accordance with the data that was determined.Table 1
Dimensions of forward speed propeller.
Location (mm)
Pitch angle (theta) (degree)
Chord length (mm)
8.71
75.51
16.22
17.42
62.67
24.08
26.13
52.22
26.97
34.84
44.06
27.26
43.55
37.75
26.35
52.26
32.83
24.96
60.97
28.94
23.46
69.68
25.82
21.98
78.39
23.27
20.59
87.10
21.16
19.32Figure 2
A typical design of unique propeller for forward speed maneuvering.Figure 3
Unique propeller for forward speed maneuvering.


## 2.1.3. Design of Unique Propeller for VTOL Maneuvering

The key design parameters are estimated with the use of mathematical equations. The diameter, pitch, chord lengths, and pitch angles of the propeller, which is aimed at carrying out VTOL operation, are the details that are discussed. The atmospheric velocity is measured to be 4 m/s, and it is assumed that the maximum rate of ascension is 50 m/s.(10)0.5×ρ×AVTOL×VVTOLspeedofUAV2−Vo2=100%oftheBWBUAV’sweight2,(11)AreaPropellerforVTOL=0.151211375m2⇒DiameterPropellerforVTOL=17.275inches,(12)PitchPropellerforVTOL=0.6×diameterPropellerforVTOL,(13)PitchPropellerforVTOL=0.6∗17.275=10.365inches,(14)CL=2×LMax.atVTOLρ×VVTOLspeedofUAV2×AVTOL.The aforementioned equation (14) yields a lift coefficient of 0.982222222. After establishing this estimate, the Reynolds number is obtained by applying the common formulas provided in Equation (7). The VTOL operation propeller lacks the NACA 25112 aerofoil, which increases the forward propeller’s coefficient of drag over the VTOL operation propeller. The chord length, pitch, and locations in the propeller’s chords for forward speed were determined using Equations (8) and (9) previously mentioned. The important relationships such as thrust and pitch of FWUAV’s VTOL propeller are expressed in Equations (10) and (12), respectively. Table 2 contains the findings of this calculation. These authors are using a coaxial propeller technology, which enables them to design propellers that revolve both clockwise and anticlockwise using CAD design software. Figures 4 and 5 display, respectively, the CAD models of the VTOL propellers.Table 2
Dimensions of VTOL propeller.
Location (mm)
Pitch angle (theta) (degree)
Chord length (mm)
21.94
62.364
30.542
43.88
43.68
34.36
65.82
32.48
31.35
87.76
25.523
27.54
109.695
20.91
24.183
131.634
17.66
21.42
153.573
15.261
19.166
175.512
13.43
17.306
197.45
11.981
15.76
219.39
10.81
14.46Figure 4
Conceptual design of unique propeller for VTOL maneuvering.Figure 5
Unique propeller for VTOL maneuvering.


## 2.1.4. Design of Primary Data of BWB UAV

The main design criteria that went into making the BWB UAV were its wingspan, length, planform area of the blended wing, and root chord. For these parameters to be found, a standard analytical method was used. Because of how important the wing loading is to the design process, we have decided that the wing loading for this BWB unmanned aerial vehicle should be 20 kg/m2.
(15)Swing=WOverallweightofBWBUAVW/SWingLoadingofBWBUAV,(16)Swing=16.773120=0.838655kg/m2.Equation (15) is used to calculate the planform area of this blended wing, and then, using precedents, the aspect ratio (AR) is assumed to be 4. The assumed AR has been imposed in Equation (17) that contains the link between wingspan and planform area of BWB UAV.
(17)AR=bwing2Swing,(18)Aspectratio=bwing20.838655⇒bwing=0.838655×4=1.832m.The typical relationship between length and wingspan of BWB-UAV is expressed in Equation (19). In most cases, the length-to-wingspan ratio of an unmanned aircraft falls around the range of 70–90%. The authors made the assumption that the design ratio (ƞ) for this work is 70%; therefore, if that is the case,
(19)L=ƞ×bwing,(20)LUAV=0.7×1.832=1.2824m.The relationship that exists between the chord length, wing area, and wing span is quite evident in Equation (21), and it can be observed rather simply.
(21)Swing=b×Cwing−root,(22)Cwing−root=0.8386551.832=0.458m.As a consequence of the analysis of the relevant literature, it was established that a taper ratio of 0.5 is the optimal number to employ in order to offer the lowest drag while maximizing lift at an angle of attack that is advantageous. As a direct consequence of this, the value of taper ratio that is chosen for this research is 0.5. The typical representation of taper ratio of BWB-UAV is expressed in the following equation:(23)Taperratioλ=CtCr.A high-wing design is envisioned for the purposes of this work. In order to compute chord at any span-wise position, the wingspan is always an essential aspect to take into consideration. With the assistance of these analytic approaches and a variety of additional supporting equations, the one-of-a-kind and highly advanced HBWB unmanned aerial vehicle can be constructed; the names of its components can be found in Figure6.Figure 6
Design of advanced HBWB UAV-typical isometric view.


## 2.2. Computational Fluid Dynamics (CFD)

The major objective of the cutting-edge and flexible CFD tool is to provide an accurate prediction of the aerodynamic pressures and forces (lift and drag) that the HBWB UAV will be subjected to in the foreseeable future. Within the domain of computational fluid dynamics, the discretization method is regarded as being of the utmost significance. In order to convert continuous functions into discrete functions, it is necessary to ensure that the distance between each point in the function remains unchanged during the transformation. If one makes the assumption that the flow’s maximum speeds are 75 m/s in the forward direction and 50 m/s when it is operating in VTOL mode, then one might be able to reach the conclusion that the flow does not compress. This immediately leads to the implementation of a solution that is based on the application of pressure. It is possible to use ANSYS Fluent to model the unmanned aerial vehicle that has been proposed in order to calculate the amount of drag it would generate while carrying out a range of maneuvers. Since the elaborate design leads to a rather high rate of turbulence formation, a traditionalk-epsilon turbulence model with quality of the second order is adopted. The working pressure is determined to be 101325 Pa in light of the facts presented here. Due to the fact that the maximum velocities were determined to be 75 and 50 m/s, respectively, the identical sets of inputs were applied for both sets of calculations. On the surface of the vehicle, there is not a problem with sliding, but free slip is present on the unmanned aerial vehicle. The wall is what governs the volume, and it has a set of dimensions that are predetermined and dictated by the geometry of a cylinder. The wingspan is ten times greater than the length of the fuselage, while the length of the fuselage is fifteen times greater than the length of the wings. Figure 7 not only gives a concise example of the border conditional components but also indicates that the control volume has the shape of a cylinder.Figure 7
Control volume of CFD problem with clear names.The SIMPLE approach is applied in order to derive the velocity and pressure coupling, and it is needed that all significant equations have a quality level of second order or higher. The flow is incapable of being compressed and keeps its density constant throughout its entirety, ultimately reaching a maximum finalized velocity of 75 m/s. Because this control volume has been determined, the pressure-based solution can now be used to create a forecast regarding the properties of the fluid. This prediction can be made because this control volume has been established. The primary fluid parameters, such as pressure and volume, are measured at each node of a control volume, and then, the values of those parameters are determined. These parameters include pressure and volume. When an unmanned aerial vehicle is equipped with a propeller, the risk that it may experience turbulence is significantly increased. As a direct consequence of this, we are going to make use of the turbulence that is present. The large amount of labour that would be required to adequately display the turbulent flow led to the selection of the second order as the best option to pursue.The development of a discretized unstructured grid on the BWB UAV is depicted in an accurate manner in Figure8. The identification of curvature and area reduction-based regions using BWB UAVs is primarily focused with the establishment of the grid. The process of precisely creating the grid is concluded when these facilities are utilized, and as a result, there is a large improvement in the possibility of achieving a trustworthy output.Figure 8
Discretized structure of the entire control volume–VTOL operation.


## 2.3. Fluid-Structure Interaction (FSI)

This imposed FSI puts a variety of materials through tests for forward speed and VTOL operations so that it can select the material that is most likely to be able to withstand the working environmental conditions that have been described. In order to manufacture UAVs, manufacturers need to use components that are primarily made of lightweight materials that also have a high load resistance. The incorporation of a variety of lightweight materials into structural calculations that are constrained to aerodynamic force imposing conditions allows this to be achieved. To make an estimate of the structural results, computations with the help of finite computational element analyses (FEA) are carried out. In this simulation, an unstructured mesh is required to be used. The fine mesh facility is utilized for the production of difficult components such as vertical and horizontal propellers. The entirety of the structure is compelled to make use of the mesh facilities, which are dependent on the curvature and proximity of the base object, i.e., UAV. After much deliberation, the final rating for the mesh quality was 0.984 out of 1. The intricate mesh of the BWB UAV is displayed in its entire splendor in Figure9.Figure 9
Discretized structure of a normal HBWB UAV.The most essential boundary conditions for this FSI analysis are the impact on the aerodynamic load, the fixed support, and the remote displacements, while the most important initial conditions for this FSI analysis are the mechanical properties of lightweight materials. As can be seen in Figure10, the aerodynamic load that is placed on the BWB UAV is completely taken into account and transferred using a one-way coupling FSI method.Figure 10
Load importation on advanced HBWB-VTOL maneuvering.The CFD method is utilized most frequently in the process of estimating and extracting the aerodynamic load that is placed on a BWB UAV. Aside from this UDL-based aerodynamic load, the remote displacement on the bother the propellers is excellently used to present the representation of rotary action of these proposed propellers, and support deformations are clearly involved in the representation of hold constrictions at the landing gear of BWB UAV.


## 3. Design, Results, and Discussions of Normal HBWB UAV

### 3.1. Conceptual Design of Normal HBWB UAV

The literature review was crucial in the development of a standard HBWB UAV since it provided information on aerofoil selection, design estimation of the center body, design computation of the blended body, and design computation of the main wing body [1–11, 13–15].


#### 3.1.1. Design of Blended Wing Body

This BWB unmanned aerial vehicle’s design and construction process may be broken down into two stages: the primary design data stage, and the secondary design data stage. Calculations have already been made for the outputs’ primary design data, which includes length, wingspan, root chord, and overall weight. The secondary data focuses on the design calculations of blended bodies as well as the selection of appropriate aerofoils for the various bodies used in BWB UAVs. The span-wise chord estimations in the blended wing body and main wing body areas of the aircraft are determined analytically with the assistance of Equations (24) to (29). The first iteration of the design was primarily based on a review of the relevant literature; the chord length at various places was figured out by applying a formula that was derived from the analysis of previously completed works. The following connections have been established with the assistance of a comprehensive review of the relevant literature.
(24)ChordlengthofextremeendofcenterbodyLengthofmainfuselageofBWB=0.667,(25)WingspanofentireofcenterbodyWingspanofentireBWB=0.15,(26)ChordlengthofextremeendofblendedwingbodyLengthofmainfuselageofBWB=0.357,(27)WingspanofentireofblendedwingbodyWingspanofentireBWB=0.45,(28)ChordlengthofextremeendofcenterbodyLengthofmainfuselageofBWB=0.1785,(29)WingspanofmainwingbodyWingspanofentireBWB=0.55.The estimated chord lengths versus span-wise positions for various subordinate bodies of BWB UAV are listed in Table3. Through these data, the BWB UAV is modeled.Table 3
Position of airfoils in the normal HBWB UAV.
Span-wise position (m)
Chord length (m)
Distance from leading edge (m)
0
1.2824
0
0.1374
0.854933333
0.19236
0.2748
0.6412
0.38472
0.4122
0.458
0.83356
0.916
0.229
0.837045


#### 3.1.2. Design of Main Wing Body

The configuration proposed for this normal HBWB UAV’s main wing is tapered wing so taper ratio, sweep angle, and span-wise chord design data need to be found out. With the help of Equations (24), (32), and (33), the aforesaid design parameters of wing are calculated and listed in Table 3.
(30)Taperratioλ=CWing−tipCwing−root⇒0.5=Ct0.458⇒CWing−tip=0.458∗0.5,(31)CWing−tip=0.229m,(32)CCr=1−21−λyb,where b is wingspan, λ taper ratio, wingspanb=1.832m, and b/2=1.832/2=0.916m.
(33)SweepangleθSweep=tan−1Cwing−root−CWing−tipWingspan/2,(34)θSweep=tan−10.916−0.4581.832⇒tan−10.25⇒14.03624347°.


#### 3.1.3. Selection of Aerofoil for Normal HBWB UAV

This conventional HBWB UAV requires an aerofoil with a high stalling angle of attack, a high coefficient of lift, a low coefficient of drag, and a highCL/CD ratio. These selection factors are involved in the process of picking the proper aerofoil. The field work was carried out with these selection criteria in mind, and as a result, eleven different aerofoils have been incorporated into the design of conventional BWB UAVs. NACA 2412, S8052, NACA 63(2)-415, NACA 64(1)-212MOD A, MH-108, MH-18, FX76-MP-120, AH 94-145, NLF(1)-1015, MH-91, and E374 are the eleven different aerofoils. Figure 11 is the representation that corresponds for a comprehensive report of the maximum angle of attack of all of the aerofoils. Figure 12 is the representation that corresponds for a comprehensive report of the maximum coefficient of lift of all of the aerofoils. Figure 13 is the representation that corresponds for a comprehensive report of the maximum coefficient of drag of all of the aerofoils. And Figure 14 is the representation that corresponds for a comprehensive report of the maximum CL/CD of all of the aerofoils.Figure 11
Comprehensive report of maximum “AoA” of various aerofoils.Figure 12
Comprehensive report of maximum coefficient of lift of various aerofoils.Figure 13
Comprehensive report of maximum coefficient of drag of various aerofoils.Figure 14
Comprehensive report of maximum “CL/CD” of various aerofoils.Only one airfoil, MH-91, was deemed to have the most desirable characteristics after being put through the initial round of testing. The computer-aided design (CAD) programme is used to create the subsequent components, and then, those portions use this airfoil as a template. The CAD models that were created as a direct result of the computations are presented in Figures15 and 16, respectively.Figure 15
Multiview of the normal HBWB UAV model.Figure 16
Parts of the normal HBWB UAV model.


### 3.2. CFD Analysis of Normal HBWB UAV

The CFD computation is carried out on a normal HBWB UAV using a pressure-based solver. The simulations are carried out in a manner in which the steady and turbulent flow is forced upon them. It has been decided that the VTOL velocity will be 100 kilometers per hour, and the forward speed velocity will be 75 meters per second. The results of the CFD analysis carried out on the normal HBWB UAV are presented in Figures17–20.Figure 17
Aerodynamic pressure variations on the normal HBWB UAV-forward speed (75 m/s).Figure 18
Velocity variations over the normal HBWB UAV-forward speed.Figure 19
Aerodynamic pressure variations on the normal HBWB UAV-VTOL (100 km/hr).Figure 20
Velocity variations over the normal HBWB UAV-VTOL (100 km/hr).


### 3.3. Grid Independence Study

In order to increase the reliability of the CFD research that was discussed earlier and improve its overall quality, an implementation of a grid convergence test-based validation has been carried out. The grid convergence test, which utilizes six distinct mesh situations, was able to discover which particular mesh situation was the most appropriate by determining which mesh situation was the most appropriate. Two distinct grid convergence studies are carried out, with one of them concentrating on the triggered velocity of the conventional HBWB UAV and the other centering its attention on the comparable stress of the sophisticated HBWB UAV. In both of these experiments, the focus is on conducting sensitivity assessments of the computational simulations that are placed on the UAV. The thorough analysis of the first grid independence test is illustrated in the figure that has been referred to as “Figure21.” As the normal HBWB UAV moved ahead, this figure demonstrates that differences in the induced velocity had an impact on the grid convergence. The mesh instance that is depicted in Figure 19 has been selected as the top converger because of the combined effects that it generates.Figure 21
Grid independence study–I (forward speed).


### 3.4. Discussions

The aerodynamic pressures that are acting on the surface of the UAV range from 2933 Pa to 3022 Pa, which is a pressure range that is extremely high for the normal HBWB UAV design of the UAV to be able to resist. It is possible that the structural failure will occur in a highly improbable manner because of the severe aerodynamic load impact. After conducting an analysis of regular HBWB UAV, the authors came to the conclusion that this type of UAV is structurally unsuitable for the blended wing body. As a result, they built advanced HBWB UAV, which they intend to deploy for targeted missions [12, 14–17].


## 3.1. Conceptual Design of Normal HBWB UAV

The literature review was crucial in the development of a standard HBWB UAV since it provided information on aerofoil selection, design estimation of the center body, design computation of the blended body, and design computation of the main wing body [1–11, 13–15].


### 3.1.1. Design of Blended Wing Body

This BWB unmanned aerial vehicle’s design and construction process may be broken down into two stages: the primary design data stage, and the secondary design data stage. Calculations have already been made for the outputs’ primary design data, which includes length, wingspan, root chord, and overall weight. The secondary data focuses on the design calculations of blended bodies as well as the selection of appropriate aerofoils for the various bodies used in BWB UAVs. The span-wise chord estimations in the blended wing body and main wing body areas of the aircraft are determined analytically with the assistance of Equations (24) to (29). The first iteration of the design was primarily based on a review of the relevant literature; the chord length at various places was figured out by applying a formula that was derived from the analysis of previously completed works. The following connections have been established with the assistance of a comprehensive review of the relevant literature.
(24)ChordlengthofextremeendofcenterbodyLengthofmainfuselageofBWB=0.667,(25)WingspanofentireofcenterbodyWingspanofentireBWB=0.15,(26)ChordlengthofextremeendofblendedwingbodyLengthofmainfuselageofBWB=0.357,(27)WingspanofentireofblendedwingbodyWingspanofentireBWB=0.45,(28)ChordlengthofextremeendofcenterbodyLengthofmainfuselageofBWB=0.1785,(29)WingspanofmainwingbodyWingspanofentireBWB=0.55.The estimated chord lengths versus span-wise positions for various subordinate bodies of BWB UAV are listed in Table3. Through these data, the BWB UAV is modeled.Table 3
Position of airfoils in the normal HBWB UAV.
Span-wise position (m)
Chord length (m)
Distance from leading edge (m)
0
1.2824
0
0.1374
0.854933333
0.19236
0.2748
0.6412
0.38472
0.4122
0.458
0.83356
0.916
0.229
0.837045


### 3.1.2. Design of Main Wing Body

The configuration proposed for this normal HBWB UAV’s main wing is tapered wing so taper ratio, sweep angle, and span-wise chord design data need to be found out. With the help of Equations (24), (32), and (33), the aforesaid design parameters of wing are calculated and listed in Table 3.
(30)Taperratioλ=CWing−tipCwing−root⇒0.5=Ct0.458⇒CWing−tip=0.458∗0.5,(31)CWing−tip=0.229m,(32)CCr=1−21−λyb,where b is wingspan, λ taper ratio, wingspanb=1.832m, and b/2=1.832/2=0.916m.
(33)SweepangleθSweep=tan−1Cwing−root−CWing−tipWingspan/2,(34)θSweep=tan−10.916−0.4581.832⇒tan−10.25⇒14.03624347°.


### 3.1.3. Selection of Aerofoil for Normal HBWB UAV

This conventional HBWB UAV requires an aerofoil with a high stalling angle of attack, a high coefficient of lift, a low coefficient of drag, and a highCL/CD ratio. These selection factors are involved in the process of picking the proper aerofoil. The field work was carried out with these selection criteria in mind, and as a result, eleven different aerofoils have been incorporated into the design of conventional BWB UAVs. NACA 2412, S8052, NACA 63(2)-415, NACA 64(1)-212MOD A, MH-108, MH-18, FX76-MP-120, AH 94-145, NLF(1)-1015, MH-91, and E374 are the eleven different aerofoils. Figure 11 is the representation that corresponds for a comprehensive report of the maximum angle of attack of all of the aerofoils. Figure 12 is the representation that corresponds for a comprehensive report of the maximum coefficient of lift of all of the aerofoils. Figure 13 is the representation that corresponds for a comprehensive report of the maximum coefficient of drag of all of the aerofoils. And Figure 14 is the representation that corresponds for a comprehensive report of the maximum CL/CD of all of the aerofoils.Figure 11
Comprehensive report of maximum “AoA” of various aerofoils.Figure 12
Comprehensive report of maximum coefficient of lift of various aerofoils.Figure 13
Comprehensive report of maximum coefficient of drag of various aerofoils.Figure 14
Comprehensive report of maximum “CL/CD” of various aerofoils.Only one airfoil, MH-91, was deemed to have the most desirable characteristics after being put through the initial round of testing. The computer-aided design (CAD) programme is used to create the subsequent components, and then, those portions use this airfoil as a template. The CAD models that were created as a direct result of the computations are presented in Figures15 and 16, respectively.Figure 15
Multiview of the normal HBWB UAV model.Figure 16
Parts of the normal HBWB UAV model.


## 3.1.1. Design of Blended Wing Body

This BWB unmanned aerial vehicle’s design and construction process may be broken down into two stages: the primary design data stage, and the secondary design data stage. Calculations have already been made for the outputs’ primary design data, which includes length, wingspan, root chord, and overall weight. The secondary data focuses on the design calculations of blended bodies as well as the selection of appropriate aerofoils for the various bodies used in BWB UAVs. The span-wise chord estimations in the blended wing body and main wing body areas of the aircraft are determined analytically with the assistance of Equations (24) to (29). The first iteration of the design was primarily based on a review of the relevant literature; the chord length at various places was figured out by applying a formula that was derived from the analysis of previously completed works. The following connections have been established with the assistance of a comprehensive review of the relevant literature.
(24)ChordlengthofextremeendofcenterbodyLengthofmainfuselageofBWB=0.667,(25)WingspanofentireofcenterbodyWingspanofentireBWB=0.15,(26)ChordlengthofextremeendofblendedwingbodyLengthofmainfuselageofBWB=0.357,(27)WingspanofentireofblendedwingbodyWingspanofentireBWB=0.45,(28)ChordlengthofextremeendofcenterbodyLengthofmainfuselageofBWB=0.1785,(29)WingspanofmainwingbodyWingspanofentireBWB=0.55.The estimated chord lengths versus span-wise positions for various subordinate bodies of BWB UAV are listed in Table3. Through these data, the BWB UAV is modeled.Table 3
Position of airfoils in the normal HBWB UAV.
Span-wise position (m)
Chord length (m)
Distance from leading edge (m)
0
1.2824
0
0.1374
0.854933333
0.19236
0.2748
0.6412
0.38472
0.4122
0.458
0.83356
0.916
0.229
0.837045


## 3.1.2. Design of Main Wing Body

The configuration proposed for this normal HBWB UAV’s main wing is tapered wing so taper ratio, sweep angle, and span-wise chord design data need to be found out. With the help of Equations (24), (32), and (33), the aforesaid design parameters of wing are calculated and listed in Table 3.
(30)Taperratioλ=CWing−tipCwing−root⇒0.5=Ct0.458⇒CWing−tip=0.458∗0.5,(31)CWing−tip=0.229m,(32)CCr=1−21−λyb,where b is wingspan, λ taper ratio, wingspanb=1.832m, and b/2=1.832/2=0.916m.
(33)SweepangleθSweep=tan−1Cwing−root−CWing−tipWingspan/2,(34)θSweep=tan−10.916−0.4581.832⇒tan−10.25⇒14.03624347°.


## 3.1.3. Selection of Aerofoil for Normal HBWB UAV

This conventional HBWB UAV requires an aerofoil with a high stalling angle of attack, a high coefficient of lift, a low coefficient of drag, and a highCL/CD ratio. These selection factors are involved in the process of picking the proper aerofoil. The field work was carried out with these selection criteria in mind, and as a result, eleven different aerofoils have been incorporated into the design of conventional BWB UAVs. NACA 2412, S8052, NACA 63(2)-415, NACA 64(1)-212MOD A, MH-108, MH-18, FX76-MP-120, AH 94-145, NLF(1)-1015, MH-91, and E374 are the eleven different aerofoils. Figure 11 is the representation that corresponds for a comprehensive report of the maximum angle of attack of all of the aerofoils. Figure 12 is the representation that corresponds for a comprehensive report of the maximum coefficient of lift of all of the aerofoils. Figure 13 is the representation that corresponds for a comprehensive report of the maximum coefficient of drag of all of the aerofoils. And Figure 14 is the representation that corresponds for a comprehensive report of the maximum CL/CD of all of the aerofoils.Figure 11
Comprehensive report of maximum “AoA” of various aerofoils.Figure 12
Comprehensive report of maximum coefficient of lift of various aerofoils.Figure 13
Comprehensive report of maximum coefficient of drag of various aerofoils.Figure 14
Comprehensive report of maximum “CL/CD” of various aerofoils.Only one airfoil, MH-91, was deemed to have the most desirable characteristics after being put through the initial round of testing. The computer-aided design (CAD) programme is used to create the subsequent components, and then, those portions use this airfoil as a template. The CAD models that were created as a direct result of the computations are presented in Figures15 and 16, respectively.Figure 15
Multiview of the normal HBWB UAV model.Figure 16
Parts of the normal HBWB UAV model.


## 3.2. CFD Analysis of Normal HBWB UAV

The CFD computation is carried out on a normal HBWB UAV using a pressure-based solver. The simulations are carried out in a manner in which the steady and turbulent flow is forced upon them. It has been decided that the VTOL velocity will be 100 kilometers per hour, and the forward speed velocity will be 75 meters per second. The results of the CFD analysis carried out on the normal HBWB UAV are presented in Figures17–20.Figure 17
Aerodynamic pressure variations on the normal HBWB UAV-forward speed (75 m/s).Figure 18
Velocity variations over the normal HBWB UAV-forward speed.Figure 19
Aerodynamic pressure variations on the normal HBWB UAV-VTOL (100 km/hr).Figure 20
Velocity variations over the normal HBWB UAV-VTOL (100 km/hr).


## 3.3. Grid Independence Study

In order to increase the reliability of the CFD research that was discussed earlier and improve its overall quality, an implementation of a grid convergence test-based validation has been carried out. The grid convergence test, which utilizes six distinct mesh situations, was able to discover which particular mesh situation was the most appropriate by determining which mesh situation was the most appropriate. Two distinct grid convergence studies are carried out, with one of them concentrating on the triggered velocity of the conventional HBWB UAV and the other centering its attention on the comparable stress of the sophisticated HBWB UAV. In both of these experiments, the focus is on conducting sensitivity assessments of the computational simulations that are placed on the UAV. The thorough analysis of the first grid independence test is illustrated in the figure that has been referred to as “Figure21.” As the normal HBWB UAV moved ahead, this figure demonstrates that differences in the induced velocity had an impact on the grid convergence. The mesh instance that is depicted in Figure 19 has been selected as the top converger because of the combined effects that it generates.Figure 21
Grid independence study–I (forward speed).


## 3.4. Discussions

The aerodynamic pressures that are acting on the surface of the UAV range from 2933 Pa to 3022 Pa, which is a pressure range that is extremely high for the normal HBWB UAV design of the UAV to be able to resist. It is possible that the structural failure will occur in a highly improbable manner because of the severe aerodynamic load impact. After conducting an analysis of regular HBWB UAV, the authors came to the conclusion that this type of UAV is structurally unsuitable for the blended wing body. As a result, they built advanced HBWB UAV, which they intend to deploy for targeted missions [12, 14–17].


## 4. Design, Results, and Discussions of Advanced HBWB UAV

### 4.1. Conceptual Design

In this work, a new way of constructing HBWB UAVs with a blended body and main wing body is suggested. This kind of HBWB UAV is called an “advanced HBWB UAV.” Based on the literature review and the construction of a normal HBWB UAV, this new design proposal for a BWB UAV is suggested for real-time applications. Table4 shows in detail the relationships between the parts that went into making this advanced HBWB UAV.Table 4
Unique relationships involved in this advanced HBWB UAV.
Sl. No.
Design data in wingspan-wise
Design data in chord length wise
Span position (m)
Relationship factor (span)
Chord length (m)
Relationship factor (chord)
1
0 (root)
0
1.2824
1
2
0.074425
0.8125
1.15416
0.9
3
0.14885
0.1625
1.02592
0.8
4
0.2977
0.325
0.89768
0.7
5
0.37785
0.4125
0.67784
0.5285
6
0.458
0.50
0.458
0.3571
7
0.916
1
0.229
0.1786


#### 4.1.1. Design of Entire Advanced HBWB UAV

The center body, the blended body, the main wing, the horizontal stabilizer, and the vertical stabilizer make up the entirety of the components that make up an advanced HBWB UAV. In addition to all of the other parts, the horizontal stabilizer was the one that played the most important role in the assembly of the advanced HBWB UAV. The presence of a coaxial propulsion system for VTOL operation causes the blended body to be longer than the center body. After the center body, the trailing and leading edges of the blended and wing bodies are assembled in a nonuniform way. This is because, up to the center body, all of the trailing edges are assembled in a single line, which forms the place for the horizontal stabilizer. The following are the novel connections that have been suggested for the development of an advanced HBWB UAV:(35)ChordlengthofextremeendofcenterbodylengthofmainfuselageofBWB=0.65,WingspanofentireofcenterbodyWingspanofentireBWB=0.30,ChordlengthofextremeendofblendedwingbodyLengthofmainfuselageofBWB=0.357,WingspanofentireofblendedwingbodyWingspanofentireBWB=0.20,ChordlengthofextremeendofcenterbodylengthofmainfuselageofBWB=0.1785,WingspanofmainwingbodyWingspanofentireBWB=0.50.


#### 4.1.2. Design of Vertical Tail and Horizontal Tail

To estimate the design data of vertical tail, the standard as well as optimized Equations (36), (37), and (38) are used.
(36)WSV−Tail=WV−StablizerSV−Tail⇒WV−Stablizer×ARV−TailbV−Tail2,(37)ARV−Tail=bV−Tail2SV−Tail=bV−TailCV−Tail−root,(38)WV−Stablizer=0.44∗Wo2/3⇒0.44×16.77312/3=2.884kg,Thus,Wo/WV−Stablizer=16.7731/2.884=5.82.In this work, the aspect ratio is assumed as 2,ARV−Tail=2.
(39)ARV−Tail=bV−Tail2SV−Tail=bV−TailCV−Tail−root,ARV−Tail=bV−TailCV−Tail−root⇒bV−Tail=2×0.229⇒0.458m,WSV−Tail=WV−StablizerSV−Tail⇒WV−Stablizer×ARV−TailbV−Tail2=2.884×20.458×0.458=27.5kg/m2.For this work, the author finalized to fix two vertical stabilizers at the end of both wind tips, thus the finalbV−Tail=0.229m and W/SV−Tail=13.75kg/m2.
(40)λV−Tail=0.3,λV−Tail=CV−Tail−tipCV−Tail−root,λV−Tail=CV−Tail−tipCV−Tail−root⇒CV−Tail−tip=0.3×0.229=0.0687m,θSweep−V−Tail=tan−1CV−Tail−root−CV−Tail−tipHalftailspanofverticaltail⇒tan−10.229−0.06870.229⇒tan−10.7⇒35°.The standard as well as advanced analytical relationships involved in the horizontal tail are given below:(41)WV−Stablizer=0.44×Wo2/3=0.44×16.77312/3=2.884kg.Thus,Wo/WV−Stablizer=16.7731/2.884=5.82.In this work, the aspect ratio is assumed as 2,ARV−Tail=2.60.
(42)ARV−Tail=bV−Tail2SV−Tail=bV−TailCV−Tail−root,ARV−Tail=bV−TailCV−Tail−root⇒bV−Tail=2.60×0.229⇒0.5954m.


#### 4.1.3. Estimation of Coefficient of Lift for Wing

At steady level flight, the important aerodynamic force relationships arelift=weightandthrust=drag. The important thrust estimation formula is given in the following equation:
(43)T=VForward22×ρ×CL×Swing⇒16.7731×9.81×2=50×50×1.2256×CL×0.838655,CL=16.7731×9.81×250×50×1.2256×0.838655=329.0882222569.63892⇒CL=0.1281.After the execution of judgment on the normal HBWB UAV, the advanced cum unique design adjustments are carried out, and the advanced HBWB UAV is created as a result with the assistance of conventional formulas. Tables5 and 6 include the final design data, in which the aerofoils for the wing and the stabilizers are chosen based on having a low coefficient of drag value, as can be seen in Figure 22. This decision was made based on the data shown in those tables.Table 5
Conceptual Design of wing of BWB-UAV.
Sl. No.
Span position (m)
Chord length (m)
Aerofoil
Swept angle
1
0 [root]
1.2824
NACA 1408
Not applicable
2
0.074425
1.15416
NACA 1408
Not applicable
3
0.14885
1.02592
NACA 1408
Not applicable
4
0.2977
0.89768
NACA 1408
Not applicable
5
0.37785
0.67784
NACA 1408
Not applicable
6
0.458
0.458
NACA 1408
14.04 degrees
7
0.916
0.229
NACA 1408
Not applicableTable 6
Conceptual design of inverted vertical stabilizer cum landing gear.
Sl. No.
Span position
Chord length
Aerofoil
Swept angle
1
0.916
0.229
NACA 0010
Not applicable
2
0.916
0.229
NACA 0010
Not applicable
3
Near to the nose
0.229
NACA 0010
Not applicable
The vertical stabilizer is planned to locate without a taper.Figure 22
Comprehensive Investigations on the aerofoil section for wing.Figure22 shows that the NACA 1408 produced a low CD when compared to other aerofoils under the operating conditions of this mission. As a result, the same aerofoil is being considered for involvement in the building of the wing of the advanced HBWB UAV. It was found that MH-91 is the best aerofoil for the implementation of the BWB UAV in the initial version, but NACA 1408 was selected as the best aerofoil for the advanced HBWB UAV since it produces three times less drag than MH-91. Therefore, the advanced HBWB unmanned aerial vehicle that has been proposed is the most aerodynamically sound option.Through the help of the above calculated procedures and its outcomes, the conceptual design of the advanced HBWB UAV is modeled, which is revealed in Figures23 and 24.Figure 23
Multiview of the second BWB model.Figure 24
Parts of the advanced HBWB UAV.


### 4.2. Results: CFD

The analysis using computational fluid dynamics (CFD) is carried out based on the boundary conditions that were discussed in the introduction, and the results are depicted in Figures25–28. The advanced HBWB unmanned aerial vehicle’s aerodynamic pressure is shown to change in response to variations in forward speed and VTOL maneuvering in Figures 25 and 27, respectively. Positive pressures are related to an increase in static pressure brought on by the movement of the UAV, whereas negative pressures are related to an increase in dynamic pressure brought on by the movement of the UAV. Both of these pressures are brought about by the movement of the UAV. Both of these shifts in pressure are the direct result of the movement of the UAV. The velocities that are distributed over the advanced HBWB UAV are depicted in Figures 26 and 28, respectively, when the UAV is moving forward and when it is performing VTOL maneuvers. When determining the operating parameters for the components of an advanced HBWB UAV, each and every one of the aerodynamic forces, including drag, lift, and side forces, have been accurately modeled. This is the case for both forward speed and VTOL maneuvering. Understanding how the UAV will respond to its surroundings and ultimately triumph over itself by utilizing its propeller and the associated rotational speeds is the primary objective of the drag force analysis. In addition, the calculation of lift will make it possible for the unmanned aerial vehicle to react and then carry out the hovering by adjusting the rotational rates of its propeller [18–34].Figure 25
Variations in the amount of pressure exerted on the UAV while travelling forward.Figure 26
Distributions of velocities across the UAV in the case of forward speed.Figure 27
Variations in pressure experienced by the UAV during VTOL flight case.Figure 28
Variations in the speed of the UAV while it is performing VTOL maneuvers.


### 4.3. FSI Results: VTOL

The loads are then applied to the outer surface of the UAV once the results of the aerodynamic simulation have been extracted. Both forward speed and vertical takeoff and landing (VTOL) operations are explored while subjected to the most extreme aerodynamic loads. The exhaustive computations are carried out in accordance with the boundary requirements that were outlined earlier and the grid convergence result that was decided upon. Figures29–32 illustrate the outcomes of the structural analysis that were conducted. In light of the fact that a vertical takeoff and landing is such a complex maneuver, additional caution has been exercised throughout this inquiry. This is due to the fact that the structural implications of VTOL are far greater than those caused by other movements. When compared to the performance of other lightweight materials, the composite made out of Carbon Fiber-Reinforced Polymer- (CFRP-) Woven-Wet and the lightweight material made out of Glass Fiber-Reinforced Polymer- (GFRP-) S-UD have superior results. Figures 29 and 30 demonstrate the structural outcomes of using CFRP, whereas Figures 31 and 32 demonstrate the structural outcomes of using GFRP.Figure 29
Variations of equivalent stress on CFRP-UAV.Figure 30
Distributions of total deformation of CFRP-UAV.Figure 31
Variations of equivalent stress on GFRP-UAV.Figure 32
Distributions of total deformation of GFRP-UAV.


#### 4.3.1. CFRP-Woven-Wet

(1) GFRP-S-UD. The comparative analysis that was carried out during the second iteration of the grid convergence test is depicted graphically in Figure 33. The changes in equivalent stress that were brought about by the movement of a sophisticated HBWB UAV in the vertical direction were what defined the point at which the grid would converge. The mesh displayed as example 3 in Figure 33 has been selected as the best converger because of the cumulative impacts it has.Figure 33
Grid independent study–II (VTOL–FSI).


### 4.4. Comparative Discussions

The comprehensive structural results of VTOL are shown in Figures34 and 35, and the lightweight materials that are imposed on this VTOL structural calculation are as follows: The various types of composites that can be utilized are as follows: aluminum alloy, CFRP-UD-Prepreg, CFRP-Woven-Prepreg, CFRP-Woven-Wet, GFRP-FR-4-Woven, and GFRP-S-UD. Aluminum alloy is also an option.Figure 34
Comparative equivalent stress outcomes under VTOL maneuvering.Figure 35
Comparative deformation outcomes under VTOL maneuvering.From Figures34 and 35, GFRP-S-UD is reacted lower in deformation and equivalent stress than all other lightweight materials so the same GFRP-S-UD is picked as the best one to resist VTOL-based aerodynamic loadings.


### 4.5. FSI Results: Forward Speed

After the VTOL operation has been performed without issue, the forward speed operation is computed, with the primary affected component of the UAV travelling at a bit quicker forward speed than the rest of the UAV. This is done so that the forward speed operation can be completed successfully. The compartmentalized architecture of an UAV is depicted in Figure36.While the associated stress and total deformation are being tested on the UAV, the material property of GFRP-Woven is being used. The results of these tests may be seen in Figures 37 and 38.Figure 36
Mesh of advanced HBWB UAV.Figure 37
Variations of equivalent stress on GFRP-UAV.Figure 38
Distributions of total deformation of GFRP-UAV.The transfer of the aerodynamic load was accomplished by the application of the FSI coupling approach, and as a direct consequence of this, the UDL was applied to the surface of the UAV. The root section of each of the blades on the propeller is equipped with the fixed support that is given. This investigation will concentrate almost entirely on material that is significantly lightweight and the mechanical features of that material. A comprehensive statement of all of the entire boundary conditions may be found in Figure39.Figure 39
Aerodynamic loads imported for the advanced HBWB UAV–forward.


#### 4.5.1. FR-4-GFRP-Woven

When the UAV is flying forward at high speeds, the landing gear and the propeller roots are the parts that experience the largest amount of stress and deformation as a result of the force exerted by the flight. Figures40 and 41, respectively, present the stresses and deformations that correlate to a number of the most cutting-edge lightweight as well as real-time implementing materials. These materials are shown to have relatively low levels of both stress and deformation. Materials with certain properties, such as GFRP-FR-4-Woven, CFRP-UD-Prepreg, and polyethylene plastic, are required for high-velocity forward motion. GFRP-FR-4-Woven is chosen in place of the other two lightweight materials because it is superior in terms of its capacity to withstand forward speed-based aerodynamic loading over an extended period of time. This is the primary reason for the material’s selection.Figure 40
Comparative outcomes deformations under forward speed maneuvering.Figure 41
Comparative outcomes equivalent stress under forward speed maneuvering.


### 4.6. Discussions

When compared to other materials, the results of the forward speed FSI demonstrate that FR-4-GFRP-Woven does, in fact, have an identical stress of 1004.8 MPa and a maximum displacement of 178.31 mm. The data on forward speed reveals that FR-4-GFRP-Woven is a preferable option to other materials due to its lower levels of stress and deformation in comparison to those levels. After carrying out FSI on a wide variety of materials for use in VTOL applications, we found that GFRP-S-UD had the lowest equivalent stress, measuring 15.122 MPa, as well as the lowest total deformation, measuring 0.66801 mm. This was the case despite the fact that it measured the lowest total deformation. This was the case in spite of the fact that it exhibited the greatest overall degree of distortion. The findings of VTOL indicate that the option known as GFRP-S-UD is the one that should be selected since it is subjected to the least amount of stress and deformation in comparison to the other possibilities. After analyzing the results of both VTOL and HTOL, the designers of the UAV settled on GFRP-S-UD as the material to use for the construction of the final version of the vehicle. This was due to the fact that the material has excellent load-bearing capacities [12, 14–17]. As can be shown in Table 7, the upgraded HBWB UAV that has been proposed possesses a number of advantageous characteristics. These characteristics include a high lift generation, a low drag production, a low deformation, a low equivalent stress induction, and a high lifetime. These are just a few of the many benefits that come along with having an upgraded HBWB UAV. As a direct consequence of this fact, the authors of this paper advocate for novel methods to the design of HBWB unmanned aerial vehicles and the proposed applications of these novel approaches in the real world.Table 7
Comparative CFD and FSI results of both normal and advanced HBWB UAVs.
Vertical take-off and landing (VTOL) operation (50 m/s)
Aerodynamic force on normal HBWB UAV (N)
Aerodynamic force on normal HBWB UAV (N)
1540.25
1980.59
Total deformation (mm) of normal HBWB UAV
Total deformation (mm) of advanced HBWB UAV
Equivalent stress in normal HBWB UAV (MPa)
Equivalent stress in advanced HBWB UAV (MPa)
5.247
0.66801
32.125
15.122
Forward speed operation (75 m/s)
Drag on normal HBWB UAV (N)
Drag on advanced HBWB UAV (N)
Lift on normal HBWB UAV (N)
Lift on advanced HBWB UAV (N)
132.56
113.535
114.26
158.508
Total deformation (mm) of normal HBWB UAV
Total deformation (mm) of advanced HBWB UAV
Equivalent stress in normal HBWB UAV (MPa)
Equivalent stress in advanced HBWB UAV (MPa)
300.12
178.31
1524.221
1004.8


### 4.7. Validation Study

Since the computational investigations play a major role in this paper, the validation study is carried out. The platform of this validation work is the propeller of this UAV. Figures42 and 43 revealed the typical views of 3-D printed propeller and its experimental setup. Figure 44 reveals the computational outcome of the propeller [18–37].Figure 42
Propeller 3-D model.Figure 43
Propeller experimental model.Figure 44
Computational outcome of the propeller with rotodynamic loading conditions.All the obtained results are comprehensively listed in Table8. From Table 8, it is observed that the imposed computational procedures are having the full capacity to provide reliable outcomes. Therefore, the same computational procedures are used in all other computational cases.Table 8
Comprehensive experimental outcomes of thrust of various propellers.
Throttle
Experimental outcome
Computational outcome
Thrust (grams)
Thrust (grams)
25%
22
27
50%
78
83
75%
115
120
100%
200
210


## 4.1. Conceptual Design

In this work, a new way of constructing HBWB UAVs with a blended body and main wing body is suggested. This kind of HBWB UAV is called an “advanced HBWB UAV.” Based on the literature review and the construction of a normal HBWB UAV, this new design proposal for a BWB UAV is suggested for real-time applications. Table4 shows in detail the relationships between the parts that went into making this advanced HBWB UAV.Table 4
Unique relationships involved in this advanced HBWB UAV.
Sl. No.
Design data in wingspan-wise
Design data in chord length wise
Span position (m)
Relationship factor (span)
Chord length (m)
Relationship factor (chord)
1
0 (root)
0
1.2824
1
2
0.074425
0.8125
1.15416
0.9
3
0.14885
0.1625
1.02592
0.8
4
0.2977
0.325
0.89768
0.7
5
0.37785
0.4125
0.67784
0.5285
6
0.458
0.50
0.458
0.3571
7
0.916
1
0.229
0.1786


### 4.1.1. Design of Entire Advanced HBWB UAV

The center body, the blended body, the main wing, the horizontal stabilizer, and the vertical stabilizer make up the entirety of the components that make up an advanced HBWB UAV. In addition to all of the other parts, the horizontal stabilizer was the one that played the most important role in the assembly of the advanced HBWB UAV. The presence of a coaxial propulsion system for VTOL operation causes the blended body to be longer than the center body. After the center body, the trailing and leading edges of the blended and wing bodies are assembled in a nonuniform way. This is because, up to the center body, all of the trailing edges are assembled in a single line, which forms the place for the horizontal stabilizer. The following are the novel connections that have been suggested for the development of an advanced HBWB UAV:(35)ChordlengthofextremeendofcenterbodylengthofmainfuselageofBWB=0.65,WingspanofentireofcenterbodyWingspanofentireBWB=0.30,ChordlengthofextremeendofblendedwingbodyLengthofmainfuselageofBWB=0.357,WingspanofentireofblendedwingbodyWingspanofentireBWB=0.20,ChordlengthofextremeendofcenterbodylengthofmainfuselageofBWB=0.1785,WingspanofmainwingbodyWingspanofentireBWB=0.50.


### 4.1.2. Design of Vertical Tail and Horizontal Tail

To estimate the design data of vertical tail, the standard as well as optimized Equations (36), (37), and (38) are used.
(36)WSV−Tail=WV−StablizerSV−Tail⇒WV−Stablizer×ARV−TailbV−Tail2,(37)ARV−Tail=bV−Tail2SV−Tail=bV−TailCV−Tail−root,(38)WV−Stablizer=0.44∗Wo2/3⇒0.44×16.77312/3=2.884kg,Thus,Wo/WV−Stablizer=16.7731/2.884=5.82.In this work, the aspect ratio is assumed as 2,ARV−Tail=2.
(39)ARV−Tail=bV−Tail2SV−Tail=bV−TailCV−Tail−root,ARV−Tail=bV−TailCV−Tail−root⇒bV−Tail=2×0.229⇒0.458m,WSV−Tail=WV−StablizerSV−Tail⇒WV−Stablizer×ARV−TailbV−Tail2=2.884×20.458×0.458=27.5kg/m2.For this work, the author finalized to fix two vertical stabilizers at the end of both wind tips, thus the finalbV−Tail=0.229m and W/SV−Tail=13.75kg/m2.
(40)λV−Tail=0.3,λV−Tail=CV−Tail−tipCV−Tail−root,λV−Tail=CV−Tail−tipCV−Tail−root⇒CV−Tail−tip=0.3×0.229=0.0687m,θSweep−V−Tail=tan−1CV−Tail−root−CV−Tail−tipHalftailspanofverticaltail⇒tan−10.229−0.06870.229⇒tan−10.7⇒35°.The standard as well as advanced analytical relationships involved in the horizontal tail are given below:(41)WV−Stablizer=0.44×Wo2/3=0.44×16.77312/3=2.884kg.Thus,Wo/WV−Stablizer=16.7731/2.884=5.82.In this work, the aspect ratio is assumed as 2,ARV−Tail=2.60.
(42)ARV−Tail=bV−Tail2SV−Tail=bV−TailCV−Tail−root,ARV−Tail=bV−TailCV−Tail−root⇒bV−Tail=2.60×0.229⇒0.5954m.


### 4.1.3. Estimation of Coefficient of Lift for Wing

At steady level flight, the important aerodynamic force relationships arelift=weightandthrust=drag. The important thrust estimation formula is given in the following equation:
(43)T=VForward22×ρ×CL×Swing⇒16.7731×9.81×2=50×50×1.2256×CL×0.838655,CL=16.7731×9.81×250×50×1.2256×0.838655=329.0882222569.63892⇒CL=0.1281.After the execution of judgment on the normal HBWB UAV, the advanced cum unique design adjustments are carried out, and the advanced HBWB UAV is created as a result with the assistance of conventional formulas. Tables5 and 6 include the final design data, in which the aerofoils for the wing and the stabilizers are chosen based on having a low coefficient of drag value, as can be seen in Figure 22. This decision was made based on the data shown in those tables.Table 5
Conceptual Design of wing of BWB-UAV.
Sl. No.
Span position (m)
Chord length (m)
Aerofoil
Swept angle
1
0 [root]
1.2824
NACA 1408
Not applicable
2
0.074425
1.15416
NACA 1408
Not applicable
3
0.14885
1.02592
NACA 1408
Not applicable
4
0.2977
0.89768
NACA 1408
Not applicable
5
0.37785
0.67784
NACA 1408
Not applicable
6
0.458
0.458
NACA 1408
14.04 degrees
7
0.916
0.229
NACA 1408
Not applicableTable 6
Conceptual design of inverted vertical stabilizer cum landing gear.
Sl. No.
Span position
Chord length
Aerofoil
Swept angle
1
0.916
0.229
NACA 0010
Not applicable
2
0.916
0.229
NACA 0010
Not applicable
3
Near to the nose
0.229
NACA 0010
Not applicable
The vertical stabilizer is planned to locate without a taper.Figure 22
Comprehensive Investigations on the aerofoil section for wing.Figure22 shows that the NACA 1408 produced a low CD when compared to other aerofoils under the operating conditions of this mission. As a result, the same aerofoil is being considered for involvement in the building of the wing of the advanced HBWB UAV. It was found that MH-91 is the best aerofoil for the implementation of the BWB UAV in the initial version, but NACA 1408 was selected as the best aerofoil for the advanced HBWB UAV since it produces three times less drag than MH-91. Therefore, the advanced HBWB unmanned aerial vehicle that has been proposed is the most aerodynamically sound option.Through the help of the above calculated procedures and its outcomes, the conceptual design of the advanced HBWB UAV is modeled, which is revealed in Figures23 and 24.Figure 23
Multiview of the second BWB model.Figure 24
Parts of the advanced HBWB UAV.


## 4.1.1. Design of Entire Advanced HBWB UAV

The center body, the blended body, the main wing, the horizontal stabilizer, and the vertical stabilizer make up the entirety of the components that make up an advanced HBWB UAV. In addition to all of the other parts, the horizontal stabilizer was the one that played the most important role in the assembly of the advanced HBWB UAV. The presence of a coaxial propulsion system for VTOL operation causes the blended body to be longer than the center body. After the center body, the trailing and leading edges of the blended and wing bodies are assembled in a nonuniform way. This is because, up to the center body, all of the trailing edges are assembled in a single line, which forms the place for the horizontal stabilizer. The following are the novel connections that have been suggested for the development of an advanced HBWB UAV:(35)ChordlengthofextremeendofcenterbodylengthofmainfuselageofBWB=0.65,WingspanofentireofcenterbodyWingspanofentireBWB=0.30,ChordlengthofextremeendofblendedwingbodyLengthofmainfuselageofBWB=0.357,WingspanofentireofblendedwingbodyWingspanofentireBWB=0.20,ChordlengthofextremeendofcenterbodylengthofmainfuselageofBWB=0.1785,WingspanofmainwingbodyWingspanofentireBWB=0.50.


## 4.1.2. Design of Vertical Tail and Horizontal Tail

To estimate the design data of vertical tail, the standard as well as optimized Equations (36), (37), and (38) are used.
(36)WSV−Tail=WV−StablizerSV−Tail⇒WV−Stablizer×ARV−TailbV−Tail2,(37)ARV−Tail=bV−Tail2SV−Tail=bV−TailCV−Tail−root,(38)WV−Stablizer=0.44∗Wo2/3⇒0.44×16.77312/3=2.884kg,Thus,Wo/WV−Stablizer=16.7731/2.884=5.82.In this work, the aspect ratio is assumed as 2,ARV−Tail=2.
(39)ARV−Tail=bV−Tail2SV−Tail=bV−TailCV−Tail−root,ARV−Tail=bV−TailCV−Tail−root⇒bV−Tail=2×0.229⇒0.458m,WSV−Tail=WV−StablizerSV−Tail⇒WV−Stablizer×ARV−TailbV−Tail2=2.884×20.458×0.458=27.5kg/m2.For this work, the author finalized to fix two vertical stabilizers at the end of both wind tips, thus the finalbV−Tail=0.229m and W/SV−Tail=13.75kg/m2.
(40)λV−Tail=0.3,λV−Tail=CV−Tail−tipCV−Tail−root,λV−Tail=CV−Tail−tipCV−Tail−root⇒CV−Tail−tip=0.3×0.229=0.0687m,θSweep−V−Tail=tan−1CV−Tail−root−CV−Tail−tipHalftailspanofverticaltail⇒tan−10.229−0.06870.229⇒tan−10.7⇒35°.The standard as well as advanced analytical relationships involved in the horizontal tail are given below:(41)WV−Stablizer=0.44×Wo2/3=0.44×16.77312/3=2.884kg.Thus,Wo/WV−Stablizer=16.7731/2.884=5.82.In this work, the aspect ratio is assumed as 2,ARV−Tail=2.60.
(42)ARV−Tail=bV−Tail2SV−Tail=bV−TailCV−Tail−root,ARV−Tail=bV−TailCV−Tail−root⇒bV−Tail=2.60×0.229⇒0.5954m.


## 4.1.3. Estimation of Coefficient of Lift for Wing

At steady level flight, the important aerodynamic force relationships arelift=weightandthrust=drag. The important thrust estimation formula is given in the following equation:
(43)T=VForward22×ρ×CL×Swing⇒16.7731×9.81×2=50×50×1.2256×CL×0.838655,CL=16.7731×9.81×250×50×1.2256×0.838655=329.0882222569.63892⇒CL=0.1281.After the execution of judgment on the normal HBWB UAV, the advanced cum unique design adjustments are carried out, and the advanced HBWB UAV is created as a result with the assistance of conventional formulas. Tables5 and 6 include the final design data, in which the aerofoils for the wing and the stabilizers are chosen based on having a low coefficient of drag value, as can be seen in Figure 22. This decision was made based on the data shown in those tables.Table 5
Conceptual Design of wing of BWB-UAV.
Sl. No.
Span position (m)
Chord length (m)
Aerofoil
Swept angle
1
0 [root]
1.2824
NACA 1408
Not applicable
2
0.074425
1.15416
NACA 1408
Not applicable
3
0.14885
1.02592
NACA 1408
Not applicable
4
0.2977
0.89768
NACA 1408
Not applicable
5
0.37785
0.67784
NACA 1408
Not applicable
6
0.458
0.458
NACA 1408
14.04 degrees
7
0.916
0.229
NACA 1408
Not applicableTable 6
Conceptual design of inverted vertical stabilizer cum landing gear.
Sl. No.
Span position
Chord length
Aerofoil
Swept angle
1
0.916
0.229
NACA 0010
Not applicable
2
0.916
0.229
NACA 0010
Not applicable
3
Near to the nose
0.229
NACA 0010
Not applicable
The vertical stabilizer is planned to locate without a taper.Figure 22
Comprehensive Investigations on the aerofoil section for wing.Figure22 shows that the NACA 1408 produced a low CD when compared to other aerofoils under the operating conditions of this mission. As a result, the same aerofoil is being considered for involvement in the building of the wing of the advanced HBWB UAV. It was found that MH-91 is the best aerofoil for the implementation of the BWB UAV in the initial version, but NACA 1408 was selected as the best aerofoil for the advanced HBWB UAV since it produces three times less drag than MH-91. Therefore, the advanced HBWB unmanned aerial vehicle that has been proposed is the most aerodynamically sound option.Through the help of the above calculated procedures and its outcomes, the conceptual design of the advanced HBWB UAV is modeled, which is revealed in Figures23 and 24.Figure 23
Multiview of the second BWB model.Figure 24
Parts of the advanced HBWB UAV.


## 4.2. Results: CFD

The analysis using computational fluid dynamics (CFD) is carried out based on the boundary conditions that were discussed in the introduction, and the results are depicted in Figures25–28. The advanced HBWB unmanned aerial vehicle’s aerodynamic pressure is shown to change in response to variations in forward speed and VTOL maneuvering in Figures 25 and 27, respectively. Positive pressures are related to an increase in static pressure brought on by the movement of the UAV, whereas negative pressures are related to an increase in dynamic pressure brought on by the movement of the UAV. Both of these pressures are brought about by the movement of the UAV. Both of these shifts in pressure are the direct result of the movement of the UAV. The velocities that are distributed over the advanced HBWB UAV are depicted in Figures 26 and 28, respectively, when the UAV is moving forward and when it is performing VTOL maneuvers. When determining the operating parameters for the components of an advanced HBWB UAV, each and every one of the aerodynamic forces, including drag, lift, and side forces, have been accurately modeled. This is the case for both forward speed and VTOL maneuvering. Understanding how the UAV will respond to its surroundings and ultimately triumph over itself by utilizing its propeller and the associated rotational speeds is the primary objective of the drag force analysis. In addition, the calculation of lift will make it possible for the unmanned aerial vehicle to react and then carry out the hovering by adjusting the rotational rates of its propeller [18–34].Figure 25
Variations in the amount of pressure exerted on the UAV while travelling forward.Figure 26
Distributions of velocities across the UAV in the case of forward speed.Figure 27
Variations in pressure experienced by the UAV during VTOL flight case.Figure 28
Variations in the speed of the UAV while it is performing VTOL maneuvers.


## 4.3. FSI Results: VTOL

The loads are then applied to the outer surface of the UAV once the results of the aerodynamic simulation have been extracted. Both forward speed and vertical takeoff and landing (VTOL) operations are explored while subjected to the most extreme aerodynamic loads. The exhaustive computations are carried out in accordance with the boundary requirements that were outlined earlier and the grid convergence result that was decided upon. Figures29–32 illustrate the outcomes of the structural analysis that were conducted. In light of the fact that a vertical takeoff and landing is such a complex maneuver, additional caution has been exercised throughout this inquiry. This is due to the fact that the structural implications of VTOL are far greater than those caused by other movements. When compared to the performance of other lightweight materials, the composite made out of Carbon Fiber-Reinforced Polymer- (CFRP-) Woven-Wet and the lightweight material made out of Glass Fiber-Reinforced Polymer- (GFRP-) S-UD have superior results. Figures 29 and 30 demonstrate the structural outcomes of using CFRP, whereas Figures 31 and 32 demonstrate the structural outcomes of using GFRP.Figure 29
Variations of equivalent stress on CFRP-UAV.Figure 30
Distributions of total deformation of CFRP-UAV.Figure 31
Variations of equivalent stress on GFRP-UAV.Figure 32
Distributions of total deformation of GFRP-UAV.


### 4.3.1. CFRP-Woven-Wet

(1) GFRP-S-UD. The comparative analysis that was carried out during the second iteration of the grid convergence test is depicted graphically in Figure 33. The changes in equivalent stress that were brought about by the movement of a sophisticated HBWB UAV in the vertical direction were what defined the point at which the grid would converge. The mesh displayed as example 3 in Figure 33 has been selected as the best converger because of the cumulative impacts it has.Figure 33
Grid independent study–II (VTOL–FSI).


## 4.3.1. CFRP-Woven-Wet

(1) GFRP-S-UD. The comparative analysis that was carried out during the second iteration of the grid convergence test is depicted graphically in Figure 33. The changes in equivalent stress that were brought about by the movement of a sophisticated HBWB UAV in the vertical direction were what defined the point at which the grid would converge. The mesh displayed as example 3 in Figure 33 has been selected as the best converger because of the cumulative impacts it has.Figure 33
Grid independent study–II (VTOL–FSI).


## 4.4. Comparative Discussions

The comprehensive structural results of VTOL are shown in Figures34 and 35, and the lightweight materials that are imposed on this VTOL structural calculation are as follows: The various types of composites that can be utilized are as follows: aluminum alloy, CFRP-UD-Prepreg, CFRP-Woven-Prepreg, CFRP-Woven-Wet, GFRP-FR-4-Woven, and GFRP-S-UD. Aluminum alloy is also an option.Figure 34
Comparative equivalent stress outcomes under VTOL maneuvering.Figure 35
Comparative deformation outcomes under VTOL maneuvering.From Figures34 and 35, GFRP-S-UD is reacted lower in deformation and equivalent stress than all other lightweight materials so the same GFRP-S-UD is picked as the best one to resist VTOL-based aerodynamic loadings.


## 4.5. FSI Results: Forward Speed

After the VTOL operation has been performed without issue, the forward speed operation is computed, with the primary affected component of the UAV travelling at a bit quicker forward speed than the rest of the UAV. This is done so that the forward speed operation can be completed successfully. The compartmentalized architecture of an UAV is depicted in Figure36.While the associated stress and total deformation are being tested on the UAV, the material property of GFRP-Woven is being used. The results of these tests may be seen in Figures 37 and 38.Figure 36
Mesh of advanced HBWB UAV.Figure 37
Variations of equivalent stress on GFRP-UAV.Figure 38
Distributions of total deformation of GFRP-UAV.The transfer of the aerodynamic load was accomplished by the application of the FSI coupling approach, and as a direct consequence of this, the UDL was applied to the surface of the UAV. The root section of each of the blades on the propeller is equipped with the fixed support that is given. This investigation will concentrate almost entirely on material that is significantly lightweight and the mechanical features of that material. A comprehensive statement of all of the entire boundary conditions may be found in Figure39.Figure 39
Aerodynamic loads imported for the advanced HBWB UAV–forward.


### 4.5.1. FR-4-GFRP-Woven

When the UAV is flying forward at high speeds, the landing gear and the propeller roots are the parts that experience the largest amount of stress and deformation as a result of the force exerted by the flight. Figures40 and 41, respectively, present the stresses and deformations that correlate to a number of the most cutting-edge lightweight as well as real-time implementing materials. These materials are shown to have relatively low levels of both stress and deformation. Materials with certain properties, such as GFRP-FR-4-Woven, CFRP-UD-Prepreg, and polyethylene plastic, are required for high-velocity forward motion. GFRP-FR-4-Woven is chosen in place of the other two lightweight materials because it is superior in terms of its capacity to withstand forward speed-based aerodynamic loading over an extended period of time. This is the primary reason for the material’s selection.Figure 40
Comparative outcomes deformations under forward speed maneuvering.Figure 41
Comparative outcomes equivalent stress under forward speed maneuvering.


## 4.5.1. FR-4-GFRP-Woven

When the UAV is flying forward at high speeds, the landing gear and the propeller roots are the parts that experience the largest amount of stress and deformation as a result of the force exerted by the flight. Figures40 and 41, respectively, present the stresses and deformations that correlate to a number of the most cutting-edge lightweight as well as real-time implementing materials. These materials are shown to have relatively low levels of both stress and deformation. Materials with certain properties, such as GFRP-FR-4-Woven, CFRP-UD-Prepreg, and polyethylene plastic, are required for high-velocity forward motion. GFRP-FR-4-Woven is chosen in place of the other two lightweight materials because it is superior in terms of its capacity to withstand forward speed-based aerodynamic loading over an extended period of time. This is the primary reason for the material’s selection.Figure 40
Comparative outcomes deformations under forward speed maneuvering.Figure 41
Comparative outcomes equivalent stress under forward speed maneuvering.


## 4.6. Discussions

When compared to other materials, the results of the forward speed FSI demonstrate that FR-4-GFRP-Woven does, in fact, have an identical stress of 1004.8 MPa and a maximum displacement of 178.31 mm. The data on forward speed reveals that FR-4-GFRP-Woven is a preferable option to other materials due to its lower levels of stress and deformation in comparison to those levels. After carrying out FSI on a wide variety of materials for use in VTOL applications, we found that GFRP-S-UD had the lowest equivalent stress, measuring 15.122 MPa, as well as the lowest total deformation, measuring 0.66801 mm. This was the case despite the fact that it measured the lowest total deformation. This was the case in spite of the fact that it exhibited the greatest overall degree of distortion. The findings of VTOL indicate that the option known as GFRP-S-UD is the one that should be selected since it is subjected to the least amount of stress and deformation in comparison to the other possibilities. After analyzing the results of both VTOL and HTOL, the designers of the UAV settled on GFRP-S-UD as the material to use for the construction of the final version of the vehicle. This was due to the fact that the material has excellent load-bearing capacities [12, 14–17]. As can be shown in Table 7, the upgraded HBWB UAV that has been proposed possesses a number of advantageous characteristics. These characteristics include a high lift generation, a low drag production, a low deformation, a low equivalent stress induction, and a high lifetime. These are just a few of the many benefits that come along with having an upgraded HBWB UAV. As a direct consequence of this fact, the authors of this paper advocate for novel methods to the design of HBWB unmanned aerial vehicles and the proposed applications of these novel approaches in the real world.Table 7
Comparative CFD and FSI results of both normal and advanced HBWB UAVs.
Vertical take-off and landing (VTOL) operation (50 m/s)
Aerodynamic force on normal HBWB UAV (N)
Aerodynamic force on normal HBWB UAV (N)
1540.25
1980.59
Total deformation (mm) of normal HBWB UAV
Total deformation (mm) of advanced HBWB UAV
Equivalent stress in normal HBWB UAV (MPa)
Equivalent stress in advanced HBWB UAV (MPa)
5.247
0.66801
32.125
15.122
Forward speed operation (75 m/s)
Drag on normal HBWB UAV (N)
Drag on advanced HBWB UAV (N)
Lift on normal HBWB UAV (N)
Lift on advanced HBWB UAV (N)
132.56
113.535
114.26
158.508
Total deformation (mm) of normal HBWB UAV
Total deformation (mm) of advanced HBWB UAV
Equivalent stress in normal HBWB UAV (MPa)
Equivalent stress in advanced HBWB UAV (MPa)
300.12
178.31
1524.221
1004.8


## 4.7. Validation Study

Since the computational investigations play a major role in this paper, the validation study is carried out. The platform of this validation work is the propeller of this UAV. Figures42 and 43 revealed the typical views of 3-D printed propeller and its experimental setup. Figure 44 reveals the computational outcome of the propeller [18–37].Figure 42
Propeller 3-D model.Figure 43
Propeller experimental model.Figure 44
Computational outcome of the propeller with rotodynamic loading conditions.All the obtained results are comprehensively listed in Table8. From Table 8, it is observed that the imposed computational procedures are having the full capacity to provide reliable outcomes. Therefore, the same computational procedures are used in all other computational cases.Table 8
Comprehensive experimental outcomes of thrust of various propellers.
Throttle
Experimental outcome
Computational outcome
Thrust (grams)
Thrust (grams)
25%
22
27
50%
78
83
75%
115
120
100%
200
210


## 5. Altitude Control of Advanced HBWB UAV

For the proposed unmanned aerial vehicle with blended wings, it is important to assess the reaction of the UAV for vertical launch and for maintaining altitude. As a means of accomplishing this goal, a streamlined model for regulating the vertical movement of the vehicle is developed as an integral component of the control for the altitude system. Thez-axis acceleration is proportional to the control input for vertical movement. This relationship between the two has been investigated as a part of the design process for the altitude controller, in conjunction with the translational equation of motion. The following describes the dynamics [38] of the translational motion of the UAV:
(44)x¨=cosϕsinθcosψ+sinϕsinψUm,(45)y¨=cosϕsinθsinψ−sinϕcosψUm,(46)z¨=−g+cosϕcosθUm.The three Euler angles, roll, pitch, and yaw, are represented byϕ,θ,andψ, and x¨,y¨,andz¨ stand for the vehicle’s acceleration along the x-, y-, and z-axes. The vehicle’s mass, “m”; the acceleration of gravity, “g”; and the amount of input needed to steer the vehicle vertically, “U,” are all represented by the letters “m,” “g,” and “U,” respectively.The angular rotations of the blades are directly controlled by the control input “U,” which affects the thrust along the z-axis of the vehicle’s body. Here, we have the input U for the controller as
(47)U=F1+F2=TcΩ12+TcΩ22,where Fi denotes the thrust force generated by propeller i; Ωi is the speed of the propeller i; and Tc is the thrust constant in N/s2. For altitude control, Equation (46) is considered. While controlling the altitude, the vehicle should be maintained in a hovering state. In hovering, the roll and pitch angles of the UAV are close to zero. Therefore, the height (z-axis position) of the UAV can be described as
(48)z=∬−g+Um.Figure45 displays the Simulink model of the altitude control system. In order to determine the reference speed for the BLDC motors that directly drive the UAV during vertical takeoff, the controller first calculates the control input U from the reference height and the actual height. It then inverts Equation (47), takes the square root, and uses that value as the reference speed.Figure 45
Simulink model of the UAV’s altitude control system.When controlling the height at which an unmanned aerial vehicle flies, it is important to take into account the dynamics of the motors. In order to provide a rough approximation of the motor dynamics, we employ an empirically determined set of parameters for a low-pass filter of the first order. In this instance, the mathematical model of the motor is derived from the electrical and mechanical features of the motor. In this particular instance, the transfer function of the motor that was being simulated (propellers and all) is(49)Gmotors=3.4670.0002363s+3.533.The altitude control system is modeled in MATLAB/Simulink using a proportional-integral-derivative (PID) controller with manually adjustable gains. The PID controllers provide a simple and effective solution to control and stabilize the vehicle because they consider every variable independently within an operating range in which the vehicle’s behavior is approximately linear. Our work presents the development of a blended wing UAV’s model by considering its dynamic behavior, and based on an analysis of its involved variables, a simplification is made that makes easier the implementation of PID controllers on a single-input, single-output system. The PID controller gains are tuned manually using MATLAB/Simulink. PID gains for the tuned altitude PID controller are listed in Table9. In our simulation, the smaller PID gains are not suitable for getting the faster response without oscillations for the commanded altitude input. Because of the high gains obtained, the UAV’s response is faster and free of oscillations. We implement some constraints into our Simulink model to ensure that our simulations produce accurate results. Even though the PID gains were tuned better, the saturation block is required for limiting the motor input voltage so that the input from the controller to the motor does not exceed the maximum input voltage limit for the motor. Since the vehicle cannot go below the Earth’s surface, the integrator used to convert z-speed into z-position must have a zero lower limit. Filtering out the negative signals that would force the motor to turn in the other direction is what saturators are for [35, 39–46]. While some multirotor vehicles are capable of performing flips and even flying upside down, this is not the primary focus of our research. Figure 46 provides a more accurate portrayal of the vertical takeoff behavior of the UAV. In reality, the pilot will lift off slowly. In order to mimic this, we use a ramp input in the form of an altitude command, with the signal starting at 0 and progressively increasing with a slope of 1. In less than 5 seconds after receiving a command signal that steadily increases in altitude (as illustrated in Figure 46), the UAV is able to lift off gently and settle itself.Table 9
PID controller gains for altitude control.
Controller
Gain
Proportional
600
Integral
280
Derivative
250Figure 46
Altitude response of the UAV with the tuned PID controller in its vertical takeoff.


## 6. Conclusions

As a result of the fieldwork, the application for this task has been finalized, and it will involve efficient surveillance as well as intruder inspections in challenging places. After conducting a comprehensive analysis of the relevant literature, the payload weight was then calculated; as a consequence, the overall takeoff weight of the BWB-based UAV is currently being approximated with the assistance of historical relationships. The dimensions of the propeller, the fuselage, and the blended body are computed based on the standard formulas. The CAD model from the system design is built by using CATIA, with the estimated dimensions being supported mostly lot for the most part. In order to conceptually create the first iteration of the BWB (normal HBWB UAV), we employed historical relationships, literature surveys, and theoretical calculations. The regular HBWB UAV is subjected to flow analysis over it as well as structural analysis. Based on the findings of these analyses, the advanced HBWB UAV, which is the second iteration of the model, is constructed. This is necessary because the regular HBWB UAV is unsuitable from a structural and aerodynamic standpoint. The flow evaluations over the UAV models that were proposed have been carried out, and as a result, the findings have been created. According to the exhaustive research, it was discovered that the advanced HBWB UAV achieved superior performance in all aspects, including high lift generation, low drag production, low deformation, low equivalent stress induction, and high lifetime. Finally, the dynamics of the altitude control system for the proposed UAV have been explored using the built PID controller. It was discovered that the UAV responds more quickly in its vertical takeoff for a given altitude command. After the successful development of CFD, the FSI has been carried out on the advanced UAV for various lightweight aerospace materials, and then, GFRP-S-UD is selected as the most suitable material in order to withstand high aerodynamic loads than other materials. As a result, the authors of this work strongly suggest this new constructional approach and advanced HBWB UAV for real-time applications such as complicated environmental-based surveillances.



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*Source: 9699908-2023-02-24.xml*