Aerodynamic Design and Analysis of an Aerial Vehicle Module for Split-Type Flying Cars in Urban Transportation

Abstract

The low-altitude economy represents an important facet of emerging productive forces, and flying cars serve as key vehicles driving its development. This paper proposes an aerodynamic design for the aerial vehicle module of split-type flying cars, which meets the functional requirements for vertical takeoff, climb, and cruising, and provides a reference solution for urban air mobility. A multidisciplinary constraint-based approach was employed to define the design requirements of the aerial vehicle module, ensuring its capability to operate in various complex environments. Through theoretical analysis and Computer-Aided Design (CAD) methods, key geometric, aerodynamic, and stability parameters were developed and evaluated. After finalizing the design concept of the aerial vehicle module, aerodynamic analysis was conducted, and aerodynamic coefficients were assessed using Computational Fluid Dynamics (CFD) simulations across angles of attack ranging from −5° to 20°. The results indicated that the aerial vehicle module achieved a maximum lift-to-drag ratio of 13.40 at an angle of attack of 2°, and entered a stall condition at 13°. The aerodynamic design enhances the module’s stability under various operating conditions, thereby improving handling performance. Overall, the aerial vehicle module demonstrates favorable aerodynamic characteristics during low-altitude flight and low-speed cruising, satisfying the design requirements and constraints.

1. Introduction

The low-altitude economy is emerging as a new engine for economic growth, with vertical aerial transportation playing a key role in the future development of cities and regions [1]. This economy is a comprehensive model centered on low-altitude flight activities, integrating technologies such as unmanned aerial systems, low-altitude intelligent networks, and other innovations. It interacts with elements such as airspace and markets, driving the development of sectors including infrastructure, aircraft manufacturing, flight operations, and support services [2,3,4]. As transportation technologies continue to advance, society is gradually entering the era of low-altitude activities, with the 0–1000 m airspace becoming an increasingly explored frontier.

The low-altitude economy represents a major future development trend, offering vast growth potential in areas such as smart transportation, emergency rescue, and three-dimensional sensing. Among these, Urban Aerial Mobility (UAM) plays a critical role, serving as a primary driver behind the emergence of new transportation industries and the overall prosperity of the low-altitude economy [5].

The Urban Aerial Mobility Network (UAMN) refers to a multi-level, three-dimensional transportation system constructed within urban airspace by leveraging low-altitude resources and integrating manned/unmanned aerial vehicles, intelligent airspace management, and digital technology support [6] (as illustrated in Figure 1). UAMN not only helps alleviate ground traffic congestion but also enhances the efficiency and flexibility of urban transportation systems [7]. In China, as vehicle ownership rates have continued to rise, the convenience and comfort of ground transportation have steadily declined [8]. Developing UAMN is essential for overcoming the limitations of traditional surface transportation by enabling efficient collaboration in three-dimensional space. This advancement facilitates the rapid flow of people and goods, improves urban traffic efficiency and safety, and optimizes spatial resource allocation. It addresses urban congestion, enhances commuting efficiency, and elevates the overall quality of urban life.

The flying car serves as a technological implementation of low-altitude vehicles within UAMN. In this paper, a flying car refers to a vehicle that functions both as a ground transportation tool and an aerial vehicle. Unlike the broader concept of “flying cars” in the eVTOL (electric Vertical Takeoff and Landing) context—which typically focuses on small low-altitude aircraft [8,9]—the flying car integrates both ground mobility and flight capabilities, functioning as an amphibious transport solution. Its ground operation resembles that of conventional vehicles, and it is designed specifically to alleviate urban traffic congestion. By enabling short-distance regional air travel between cities, flying cars address the limitations of existing road congestion and constrained ground-based takeoff conditions, thereby playing a crucial role in promoting sustainable urban mobility.

In urban low-altitude environments and short-range regional transportation, no fixed clearance zones exist for aerial vehicles. The ground and airspace conditions in such settings are more complex and variable compared to traditional airports. These conditions necessitate that low-altitude vehicles possess basic taxiing and takeoff capabilities, while primarily relying on Vertical Takeoff and Landing (VTOL) technology to initiate flight. This eliminates the dependence on long runways or specific takeoff environments [10]. VTOL technology enables the shortest aerial distance between two points, thereby improving energy efficiency and reducing flight time—particularly in congested urban areas or geographically complex regions such as mountainous terrain, river valleys, or coastal zones where direct ground transportation is limited. Moreover, VTOL systems combined with relatively high cruising speeds contribute to further reductions in total travel time.

As renewable energy generation continues to increase, the carbon intensity of most power grids is expected to decrease significantly in the coming years. Since the greenhouse gas emissions of eVTOL aircraft are directly proportional to the carbon intensity of the power grid, these aircraft are anticipated to offer a distinct advantage over traditional fossil fuel-powered vehicles. A well-designed distributed electric propulsion (DEP) system can allocate up to 70% of its power to propulsion, nearly doubling the energy transfer efficiency compared to turbofan engines, which typically operate at around 40% efficiency [11]. This enables notable improvements in cruising speed and overall aircraft performance. Furthermore, distributed electric systems exhibit lower sensitivity to battery energy density, allowing smaller motors to replace larger ones without compromising total power output. This design flexibility helps avoid the technical and economic challenges associated with high-power motors [12].

Excessive vehicle weight can severely limit the range and payload capacity of flying cars. As the ground driving function relies on the chassis structure and hub drive system, while the aerial flight function depends on aerodynamics and flight control systems, the design of split-type flying cars is divided into three modules: the aerial vehicle module, the chassis module, and the payload cabin module, with the payload cabin serving as the carrier for passengers or cargo. The modular design of the flying car allows for flexible switching between ground driving and aerial flight [13], while also enabling independent development of each module. This significantly accelerates the overall design process and facilitates the modular verification of functionality under various operational conditions [14,15].

The adoption of Unmanned Aerial Vehicle (UAV) technologies [16] can accelerate the application of flying cars in the low-altitude economy. It can decouple the roles of pilot and passenger, allowing passengers to travel without the need to operate the vehicle themselves. The aerodynamic design of a scaled-down prototype of a multi-modular flying car is discussed in this paper, whose mission is to use autonomous flight to achieve fast point-to-point transport in low-altitude urban airspace. This technology can reduce transport time and financial costs [17], while avoiding the need to train flying car passengers to become qualified pilots for aerial transport.

An in-depth exploration of the operational environment and mission objectives of the aerial vehicle module (AVM) of the multi-modular flying car is provided in this paper, aiming to clarify the design requirements and objectives for its aerodynamic design. AVM is intended for unmanned cargo transport in urban air mobility applications, primarily focusing on airspace below 1000 m above ground level. According to the building height data for Beijing, the cruising altitude of flying cars should be above 500 m. Based on Uber’s proposed UAM transportation tasks [18] and the Langley Research Center’s study on passenger transport for urban air mobility missions [19], the transportation functions of AVM should include commuting, airport shuttles, point-to-point urban transportation, and services like subways. Considering the average commuting distance in major Chinese cities, the effective range of AVM should be no less than 10 km. For safety and exceptional circumstances, the cruising speed of AVM should range between 120 and 300 km per hour: it should be capable of low-speed cruising at 120 km per hour to ensure initial safety and stability, while also being able to achieve higher speeds of 150 km per hour and even up to 300 km per hour in specific mission scenarios.

The flying car is required to operate within relatively constrained urban and mountainous environments, which necessitates that the AVM possess VTOL capabilities and the ability to rapidly gain altitude over short distances. Consequently, the selection of the wing type and rotor configuration for AVM is critical. The choice of airfoil and wing design is aimed at achieving a high lift coefficient (C_L). Since the vehicle must perform VTOL operations, it requires a sufficiently powerful propulsion system. Thus, the selected propellers and power systems must meet the performance requirements for vertical takeoff, cruising, and other operational conditions. This implies that the aircraft must maintain an appropriate power-to-weight ratio (W/P) at each phase of its flight profile. Additionally, the design of AVM faces an additional challenge: ground gusts and unstable incoming airflow inevitably affect takeoff stability. Therefore, careful consideration must be given to the selection of the wing and tail configuration to mitigate potential degradation in stability and performance.

To achieve the desired flight performance of aerial vehicles, numerous studies have focused on optimizing the aerodynamic characteristics necessary for flying cars, thereby supporting the successful development of such vehicles. Jiang et al. (2023) investigated the aerodynamic design and evaluation of a ducted fan lift system for vertical takeoff and landing (VTOL) flying cars. Their research highlighted the advantages of ducted fans in improving lift efficiency, reducing noise, and enhancing safety in urban environments, demonstrating the promising potential of this system for flying cars [20]. Xu et al. (2024) performed a numerical analysis of coaxial contra-rotating propellers in eVTOL vehicles, emphasizing the superior performance of this configuration in enhancing lift-to-drag ratios and reducing induced drag. Their findings showed that coaxial contra-rotating propellers significantly improve aerodynamic efficiency and reduce energy consumption, making them well-suited for flying car applications [21]. Zhang and Barakos (2024) explored the aerodynamic performance of redundant propellers for multi-rotor flying cars during cruise flight. Their study indicated that redundant propeller configurations can enhance vehicle stability and control while reducing drag, especially in turbulent flight conditions [22]. Additionally, Zhang et al. (2024) conducted wind tunnel experiments on the nonlinear ground effect for flying cars. Their findings emphasized the substantial increase in lift and reduction in drag when flying close to the ground, which helps improve energy efficiency and stability during low-speed flight [23]. Furthermore, Magata et al. (2025) simulated forward and turning flight of flying cars, considering the impact of wind disturbances on flight dynamics, and performed a comparative evaluation of flight dynamics models. Their research highlighted the importance of optimizing aerodynamic designs to mitigate the effects of wind disturbances and ensure flight stability [24]. Together, these studies underscore the critical role of aerodynamic analysis in the development of flying cars.

In this paper, the following sections systematically describe and provide a detailed presentation of the design process and the assessments of the AVM, enabling the split-type flying car to achieve its flight capabilities. Section 2 outlines the research background related to the AVM. Section 3 presents the multi-condition constraints based on flight requirements, along with an analysis of the conceptual design’s power performance, static stability, and overall performance. Section 4 introduces the fundamental work for numerical simulations using CFD. Section 5 presents the aerodynamic coefficient calculations and performance analysis of the AVM. Section 6 discusses the flight performance of the AVM during low-altitude, low-speed flight, such as lift efficiency, induced drag, and flow field control, and compares it with other vehicles designed for similar missions. Section 7 concludes the study and discusses future research directions.

2. Research Background

To reduce the overall weight of the airframe during flight, thereby extending the endurance and improving the payload capacity of the flying car, the multi-modular split-body configuration is considered one of the most promising technological approaches currently receiving considerable attention in the industry [25]. As shown in Figure 2, in the design of dual-mode “land-air” flying cars, most automotive companies and academic institutions have adopted a multi-modular split-body structural design. However, for aerial vehicle modules responsible for flight functionality, the use of a tilt-rotor and fixed-wing combination remains relatively uncommon. The design and aerodynamic analysis of modular tilt-rotor and fixed-wing aerial vehicles remains an open research topic. Compared to multi-rotor systems, the tilt-rotor configuration reduces the number of lift-generating rotors and motors required. The tilt-rotor system can fully or partially counteract vertical loads during vertical takeoff and landing, climb, and glide descent, while also providing all necessary propulsion during cruise flight [26].

Thus, an aerodynamic layout featuring four tilt-rotors and two auxiliary rotors is proposed in this study. This configuration is optimized based on a multi-rotor and fixed-wing airframe, reducing the total number of lifting elements and leading to a reduction in the overall propulsion system weight, thereby enhancing aerodynamic efficiency. However, tilt-rotor mechanisms present challenges such as increased structural weight and more complex flight control requirements. The detailed design and performance verification of these mechanisms will be addressed in future research.

The intended operational environments for flying cars include urban areas and rugged mountain valleys, both of which represent complex three-dimensional transportation zones. The urban heat island effect and the geographical features of mountainous regions can induce free-flow disturbances, generating updrafts or crosswinds that may impact the vehicle’s stability. Furthermore, the relatively low altitudes at which the flying car is expected to operate require it to exhibit favorable performance during low-speed, low-altitude flight. To address these challenges, a high-aspect-ratio wing configuration is necessary to reduce induced drag and improve stall characteristics [31].

To resolve the aforementioned issues, a dual-battery compartment structure has been designed on both sides of the fuselage, along with a larger fixed tail support structure, to enhance the stability and low-altitude flight capabilities of the AVM. The German FW-189 aircraft demonstrated during its service that the combination of a fixed tail support structure and dual-engine compartments provided excellent flight performance and stability during low-altitude operations. Ma Y. from the Brunswick University of Technology has also validated the concept of a twin-fuselage (TF) design, showing that it achieves good energy utilization efficiency when paired with a high-aspect-ratio wing configuration [32]. To mitigate the impact of uncontrollable factors such as strong winds, the AVM is equipped with a parachute system, which serves as a vertical landing device under extreme conditions.

The overall design concept of the split-type flying car can be described as consisting of three interconnectable modules. The overall configuration represents a combination of an aerial vehicle and a ground vehicle, as shown in Figure 3a. When aerial transportation is required, the payload module disconnects from the chassis module and connects to the flight module, achieving an aerial flight posture, as shown in Figure 3b. When the flight module delivers the payload module to the designated location and ground transportation becomes necessary, the flight module disconnects and the payload module reconnects with the chassis module to form the ground driving configuration, as shown in Figure 3c.

3. Conceptual Design and Performance Analysis

The conceptual design of the tilt-rotor-fixed-wing split-type flying car was derived from the potential demands of its operational environment, followed by an in-depth study of the technologies related to its flight conditions. This design approach draws from Raymer’s work [26], the Aircraft Design Manual [33] on various fixed-wing configurations [34], as well as design recommendations for multi-rotor aircraft and helicopters [35,36].

The design process of the AVM is divided into three main stages (Figure 4). The first stage involved the selection of the conceptual design for the AVM, defining its design parameters, including operational requirements, payload weight associated with performance characteristics, and determining its mission profile. During this stage, it was crucial to perform a constraint analysis under various limitations to ensure that the designed systems could meet the mission profile requirements of the AVM. The second stage involved the preliminary estimation of aerodynamic, geometric, weight, and stability parameters for the AVM. The evaluation of aerodynamic characteristics, including speed, endurance, and VTOL capabilities, was conducted. The geometry, actual dimensions, and technical specifications of the AVM are provided in Section 3.8. The third stage involved CFD simulation analysis for the AVM to determine its aerodynamic coefficients and assess its aerodynamic performance under the mission profile requirements. This stage led to the final aerodynamic shape design of the AVM.

3.1. Design Requirements and Objectives

In this paper, the design requirements for the AVM are defined through an analysis of its anticipated functional objectives and flight modes. The AVM is required to meet both the performance and operational criteria of a fixed-wing aircraft while also possessing vertical takeoff and landing capabilities. The design specifications include the vehicle’s ability to perform vertical takeoff and landing, hover, transition from hover mode to climb mode, and ultimately achieve cruising flight. The design is intended to enable rapid, short-range transportation in low-altitude environments. Specific design requirements are outlined based on the potential application scenarios of the flying car, as presented in

Table 1. Design requirements of the AVM.

Flight Performance Requirements	Payload	Aerodynamic	Structural
Vertical take-off altitude = 10 m	Mass spectrometer *	High Lift	Aluminum alloy frame
Climbing altitude ≈ 500 m	Avionics	Low Drag	Carbon fiber skin
Cruising altitude = 500 m	Lithium battery system	Low Re	Test temperature ≈ 20 °C
Cruising distance ≥ 70 km	Optical sensors	High wing	Replaceable battery compartment
Cruising speed ≥ 34 m/s	Parachute	Tilting rotors
Payload = 50 kg	Transport load

* For flying cars used in security, firefighting, and emergency response, a mass spectrometer can detect: explosive residue gases, toxic industrial gases, and leaked flammable gases. This enables the assessment of site safety and supports decision-making during rescue operations.

.

3.2. Weight Estimations and Mission Profile

The weight estimation for the AVM was based on the basic operational conditions and structural design of the flying car. The first estimated value was the Maximum Take-Off Weight (MTOW), which was determined by the average test flight weight statistics of current flying cars in China. The takeoff weight was considered to be composed of the weight of onboard instruments ($W_p$), the empty weight ($W_e$), the propulsion system weight ($W_f$) [37], and the payload weight of the AVM ($W_c$).

The initial estimation of the Maximum Takeoff Weight (MTOW) is given by Equation (1):

$$MTOW=W_p+W_e+W_f+W_c$$

(1)

the weight of onboard instruments (W_p) included the total weight of all instruments required to control the AVM during its flight modes. This included the flight control system, atmospheric measurement devices, infrared camera systems, data storage equipment, and the parachute system for emergency recovery. The weight estimation of the electric propulsion system, particularly the battery system (this study selected a ternary lithium battery as the power source), was a critical aspect of the airframe design. The mass of each component was estimated using methods outlined in the literature [38]. Since the mass of an electric-powered vehicle does not decrease with flight time, the required power for different flight phases was calculated using a simplified method based on the vehicle’s mass, thrust, and aerodynamic performance [39]. Based on the mission profile (Figure 5) and each flight phase, the maximum energy consumption for each phase was calculated, and the battery mass was determined by the total energy demand and flight time.

The empty weight ($W_e$) of the airframe was estimated based on the total weight of the aircraft (MTOW). This method used historical data and empirical formulas to estimate the weight proportions of each component relative to the total aircraft weight.

The empty weight estimation formula is given by Equation (2):

$$W_{empty}=MTOW\times k$$

(2)

where $k$ is a constant, typically determined by the aircraft type, structural complexity, and design requirements. In this study, $k$ = 2/3, meaning that the empty weight (W_e) accounts for 66.7% of the aircraft’s maximum takeoff weight.

After iterative calculations, the first weight estimate is as follows:

MTOW = 300 kg, W_p = 16.31 kg, W_e = 200 kg, W_f = 100.02 kg, W_c = 100 kg.

3.3. Constraint Analysis and Propulsion System

This paper performed a constraint analysis for the airframe design requirements, deriving the ranges for wing loading (W/S) and power loading (W/P), which were necessary to meet the performance requirements of the airframe. When using propellers, the thrust-to-weight ratio (T/W) must be converted into a power-to-weight ratio (W/P). By selecting the highest wing loading and the lowest thrust loading within the allowable ranges, an aircraft design with the lightest weight and lowest cost was achieved while satisfying performance requirements [33].

For the constraint analysis of propeller-driven civil aircraft, typical considerations include takeoff distance, landing requirements, climb, cruise, and stall speed constraints [33]. By simplifying the constraint analysis formulas for each flight phase, a feasible design space can be formed that satisfies all constraints. The suggested feasible design range for the airframe (shown in the red box in Figure 6) minimizes the specific power and wing load while providing a safety margin (W/P = 0.252 N/W and W/S = 337.6 N/m²).

For the tilt-rotor airframe, which satisfies the above constraints and is capable of VTOL, a VTOL constraint analysis was also performed. In the VTOL constraint analysis, the airframe was simplified to a multi-rotor configuration. The main constraints for VTOL include the power requirement of the propulsion system and the disk loading constraint.

The VTOL constraint is given by Equation (3):

$$W_{battery}={\displaystyle \frac{\sum P_M·t}{\eta ·EnergyDensity}}$$

(3)

where $W_{battery}$ is battery weight, $P_M$ is average power required for vertical takeoff by each motor, $t$ is hover time during takeoff and landing, η is overall efficiency of the electric propulsion system, $EnergyDesity$ is energy density of the battery, with a value of 273.7 Wh/kg used in this study [40].

The disk loading constraint estimation is given by Equation (4):

$$BladeLosd={\displaystyle \frac{MTOW}{N·\pi {\left(\frac{D}{2}\right)}^2}}$$

(4)

where $Bladelosd$ is disk loading, $N$ is number of propellers, $D$ is propeller diameter.

The hover power of the AVM decreases as the rotor radius increases. This is because the increased rotor area enhances the lift efficiency of the tilt-rotor propellers, thereby reducing energy consumption. The rotor disk loading also decreases as the rotor radius increases. While this reduces the risk of rotor failure to some extent, excessively low rotor disk loading can result in the blades being unable to provide effective lift. Additionally, increasing the rotor radius enlarges the diameter of the rotor, leading to higher drag and adversely affecting the overall size design of the AVM. The analysis of the vertical takeoff power and rotor disk loading constraints (Figure 7) ensures that the AVM is capable of vertical takeoff, while keeping the required energy consumption within the design limits of the battery system and ensuring that the rotor disk loading remains within the available range. This allows the AVM to provide sufficient lift through the propellers and wings in all flight modes, while the electric propulsion system supplies adequate power for each flight phase.

To meet the demand for static thrust or effective thrust in various flight conditions, the thrust and thrust efficiency for three different fixed-pitch propellers (Figure 8) are analyzed in this study. Under the same lift condition, the propeller diameter is proportional to the rotor efficiency; while for the same blade diameter, the three-blade propeller exhibits similar efficiency to the two-blade propeller, with the three-blade configuration providing a more stable lift output. Therefore, a 52-inch three-blade propeller was selected for the four tilt-rotor propellers of the AVM airframe to ensure stable thrust/lift output. A 52-inch two-blade propeller was selected for the auxiliary rotors to provide additional lift support during vertical takeoff and landing, as well as climb. In cruise conditions, the propeller is set to a fixed-pitch mode, where it remains stationary along the direction of travel to reduce drag.

3.4. Aerodynamic Design

Aerodynamic analysis is crucial for the design of the AVM airframe, as it allows for the optimization of the airframe’s performance and stability parameters. The power-to-weight ratio (W/P) and wing loading (W/S) are key parameters in aircraft design. A higher thrust-to-weight ratio results in better performance, while lower wing loading enhances lift performance. According to the constraint analysis (Figure 6), the critical performance conditions indicate that the stall speed should not be less than 17 m/s. In special cases, the airframe can utilize the ground driving function of the flying car for taxiing takeoff, with the recommended takeoff runway length being less than 250 m. The wing surface area (S) is suggested to be approximately 6.525 m². These limitations led to an iterative design process for the primary aerodynamic surfaces.

3.4.1. Wing Design

eVTOLs and flying cars typically operate at flight speeds below 200 km/h [29], classifying them as low-speed aircraft. For such low-speed applications, two types of wings are commonly considered: the first type features a high aspect ratio, combined with a small sweep angle and low angle of attack; the second type involves a wing with an extremely low aspect ratio, resembling an extension of the aircraft’s body [41]. NASA’s research [42] indicates that high aspect ratio wings can significantly enhance the lift-to-drag ratio (C_L/C_D) at low speeds, thus improving energy efficiency. Straight wings and elliptical wings, when compared to other wing shapes, reduce induced drag at low Reynolds numbers. However, due to manufacturing complexity and the risk of complete stalling, elliptical wings are rarely used in practical aircraft applications [41]. Currently validated eVTOLs, such as the Joby S4 (Joby Aviation, Santa Cruz, CA, USA) and Archer Midnight(Archer Aviation, San Jose, CA, USA), adopt relatively straight, long wings with high aspect ratios, confirming the validity of this approach.

Flying cars typically operate in confined environments such as urban and mountainous regions, where space is limited. A large wingspan would hinder the integration of the flight module with the passenger cabin, while also imposing stricter airworthiness requirements and severely limiting maneuverability in narrow environments. Therefore, this study constrained the wingspan of the AVM to no more than three times the width of a standard lane (10.5 m), and designed it with a relatively straight, long “W”-shaped high-aspect-ratio wing. To ensure sufficient aerodynamic efficiency, the wing’s installation angle was set at 3°, maximizing aerodynamic performance.

Table 2. Wing parameters.

Wing Parameters	Value
Wingspan (b)	6.87 m
Root chord length (Cr)	0.963 m
Tip chord length (Ct)	0.638 m
Mean aerodynamic chord	0.802 m
Wing Area (S)	4.852 m2
Airfoil thickness t/c	0.12
Aspect Ratio (AR)	8.566
Installation angle	3
Operational Lift Coefficient	0.6–1.2
Maximum Lift Coefficient	1.5
Oswald efficiency (e)	0.75–0.8
Aerodynamic efficiency	15–20
Airfoil	NACA 63-412

presents the geometric and aerodynamic parameters of the wings.

3.4.2. Fuselage Design

In designing the AVM’s fuselage, three objectives had to be achieved: first, to provide structural support and connections for the wing aerodynamic surfaces; second, to provide adequate storage space for onboard flight control equipment, parachutes, and other components; and third, to accommodate the coupling device that connects the fuselage to the passenger/cargo cabin. Additionally, the design of the AVM’s fuselage needed to consider the overall aerodynamic characteristics of the aircraft after connection to the passenger/cargo cabin. Therefore, the initial fuselage design for the AVM was a streamlined rectangular structure with rounded edges. This design drew inspiration from the fairing structure on the top of cargo truck cabins (as shown in Figure 9) and was based on aerodynamic optimization studies of low-speed nacelles [43,44]. The goal was to minimize the boundary layer thickness on the top of the airframe and reduce the turbulent kinetic energy intensity at the rear.

3.4.3. Empennage Design

Considering the stability requirements of the AVM’s application scenarios, a fixed tail support structure was designed in this study, with the two vertical tail surfaces connected by a horizontal tailplane, referred to as a “twin-boom,” which is connected to the wing via the battery compartment. The “twin-boom” configuration effectively reduces vortex formation and flow instability around the tail and wing, enhancing stability while also significantly lowering induced drag, thereby improving aerodynamic efficiency [45]. In VTOL conditions, this design offers better maneuverability and stability, especially during transitions between hover and low-speed flight. Furthermore, compared to a single-tail design, the twin-boom structure provides greater overall stability, facilitates the installation of auxiliary tilting rotors, and more effectively distributes loads, reducing interference between the rotors and tail surfaces. This results in optimized aerodynamic performance in both static and dynamic flight conditions. [46,47], while enhancing maneuverability and stability. When designing the length of the horizontal tailplane, it was essential to consider the spacing between the horizontal tail and the wing. Placing the horizontal tail too close to the wing can cause vortex formation behind the wing, with the propeller slipstream amplifying the vortex, leading to significant increases in instantaneous acceleration and irregular distribution of the flow field at the wing’s trailing edge. Conversely, placing the horizontal tail too far from the wing reduces the coupling effect, which is crucial for maintaining aerodynamic stability. An excessive distance between the tail and the wing could also affect the center of gravity (CG) height, thus interfering with the aircraft’s takeoff and landing performance. Furthermore, a longer tail reduces the structural stiffness, complicating the overall design. These findings are consistent with previous studies: Altunkaya and Özkol (2022) highlighted that optimizing the horizontal tail’s placement is critical to minimizing aerodynamic instability and improving performance [48]. Similarly, Karatoprak (2019) emphasized that tail size and positioning have a direct impact on the aircraft’s stability and control, particularly during low-speed maneuvers and takeoff [49]. Leishman (2023) also noted that the tail’s length should be carefully balanced with the airframe to ensure optimal stability while maintaining ease of control during various flight phases [50]. Therefore, it is recommended that the length of the free horizontal tail should not be too short and should ideally match or slightly exceed the length of the airframe to maintain both aerodynamic efficiency and structural integrity. [51].

The fixed-tail structure generates minimal drag while providing sufficient downforce to ensure longitudinal stability at small angles of attack. The parameters of the fixed-tail structure are listed in

Table 3. Empennage parameters.

Empennage Parameters	Value
Empennage span (be)	3 m
Chord length (c)	0.55 m
Empennage area (Se)	1.65 m2
Aspect Ratio (AR)	5.45
Angle of incidence	0°
Airfoil	NACA 63-412
The height distance between the wing	0.29 m
Vertical tail fin height	0.9 m

.

3.4.4. Battery Compartment Design

In general, eVTOLs and drones often adopt designs in which the power batteries are housed within the fuselage and wings. However, this compact airframe design reduces the available space for batteries, which in turn affects the range performance. This study draws inspiration from twin-boom aircraft designs (such as the German FW-189 “Owl” reconnaissance plane, the BF-109 fighter, the American F-82 “Twin Mustang” fighter, and the P-38 “Lightning” fighter) by incorporating a structural connection between the midsection of the wings and the fixed tail support. This design allows for the expansion of the battery compartment to meet the required battery capacity.

Under vertical takeoff, landing, and hover docking conditions, the AVM must demonstrate excellent lateral wind resistance to ensure overall stability. Small aircraft typically increase the vertical tail area to improve crosswind resistance [52]. In this study, the battery compartment was vertically extended, forming a thick and flat shape, creating a structure akin to a vertical tail that provides enhanced resistance to crosswinds.

3.4.5. Propeller Nacelle Design

To ensure that the AVM met its intended mission profiles, a distributed electric propulsion system comprising six rotors—four tilting rotors and two auxiliary rotors—was designed. The four primary tilting rotors are capable of continuous tilt from 0° (horizontal) to 100°, as illustrated in Figure 10, enabling the vehicle to perform vertical takeoff and landing, climb, cruise, and unpowered emergency descent, in accordance with the angular requirements for these flight modes.

The main thrust propellers of the AVM, equipped with tilt functionality, are located at the wingtips. Mauro’s team discusses how this configuration allows the thrust propellers to generate counter-rotating vortices that oppose the natural wingtip vortices, thereby reducing overall vortex strength [53]. Dipan Deb’s team also demonstrates that, at low Reynolds numbers, the counter-rotating vortices produced by rotors provide greater lift enhancement than small wings [54]. This effect improves the effective span efficiency of the wings and significantly reduces overall wing drag.

The auxiliary lift rotors of the AVM are positioned at the top of the vertical stabilizers, as part of the fixed tailboom structure described in Section 3.4.3. This distributed drive layout enables the AVM to achieve a multi-rotor configuration during VTOL and hover conditions, enhancing overall stability. However, due to the requirement for the AVM to dock with the crew module, its center of gravity must remain as close as possible to the aircraft’s midsection. As a result, the auxiliary lift rotors create a large lever arm, affecting the pitch moment balance of the AVM. Additionally, at high angles of attack, flow separation and loss of effectiveness may occur at the wing and horizontal tailplane junction, leading to reduced stability and potential loss of control. To mitigate this issue, a mechanism with a servo-like function is required for overall stability control.

Considering these constraints, the authors incorporated an additional pair of auxiliary rotors, mounted on top of the battery compartments on both sides of the AVM. This configuration helps decouple the aerodynamic interference between the front and rear tiltrotors, particularly at high angles of attack, and facilitates attitude adjustment and torque balance during VTOL operations.

The aforementioned rotor placement design enables the flight control system to enhance the overall stability and controllability of the AVM across various flight conditions by coordinating the tiltrotors and auxiliary rotors.

3.5. Drag Polar Estimation

This paper focuses on the stability contributions of the main aerodynamic surfaces of the AVM, including the wing, fuselage, tailplane, and battery compartment (Figure 11). Specifically, the analysis examines how each of these aerodynamic surfaces affects the variation in the aircraft’s aerodynamic moment. A rapid estimation of the longitudinal static stability of the aircraft can be expressed by Equation (5) [55]:

$${\displaystyle \frac{d{C}_m}{d\alpha }}=-a_w·{\displaystyle \frac{x_{cg}-x_{ac}}{c}}·a_h·\left(1-{\displaystyle \frac{d\epsilon }{d\alpha }}\right)·{\displaystyle \frac{S_h}{S}}·{\displaystyle \frac{l_h}{c}}$$

(5)

where $C_m$ is the pitching moment coefficient, α is the angle of attack, $a_w$ is the slope of the main wing’s lift curve, $x_{cg}$ is the position of the center of gravity (expressed as a fraction of the wing chord length), $x_{ac}$ is the location of the aerodynamic center of the wing, $c$ is the mean aerodynamic chord of the wing, $S$ is the wing area, $S_h$ is the horizontal tail area, $l_h$ is the distance from the aerodynamic center of the wing to the horizontal tailplane, ${\displaystyle \frac{d\epsilon }{d\alpha }}$ is the slope of the tailplane’s lift curve.

The aerodynamic surface of the battery compartment contributes to enhancing the aircraft’s stability in crosswind conditions by increasing the yawing moment. The battery compartment acts as a large vertical stabilizer, generating a counteracting moment to the yaw angle, which forces the aircraft to return to its original heading when disturbed by crosswinds. The stronger the opposing yaw moment, the faster the aircraft will recover to its equilibrium state.

A conceptual analysis of longitudinal static stability was conducted, calculating the contributions from the wing, fuselage, and tailplane to maintain the AVM’s moment coefficient derivatives. Figure 12 illustrates the relationship between the pitching moment coefficient at the center of gravity and the angle of attack, showing how the main aerodynamic surfaces of the aircraft influence this relationship. As the angle of attack increases, the pitching moment coefficient generated by the wing gradually increases, indicating that the wing has a significant contribution to the pitching stability of the AVM with the increase in angle of attack. Meanwhile, with the increase in angle of attack, the pitching moment produced by the tailplane increases in the opposite direction. Its primary function is to counteract the pitching moment generated by the wing, thereby maintaining the aircraft’s stability. The overall pitching moment of the AVM is not sensitive to changes in the angle of attack, indicating that it can maintain good stability under various flight conditions, thus avoiding instability issues that may arise from large variations in the angle of attack.

3.6. Static Stability Analysis

The influence of all systems on the stability of the aircraft generally occurs through two main mechanisms. One mechanism involves altering the position of the center of gravity (CG), while the other mechanism involves changing the moment of inertia or generating aerodynamic forces. The location of the center of gravity is determined by the distribution of the aircraft’s weight and balance, making proper weight distribution crucial for ensuring the aircraft’s stability.

According to the statistical method proposed by Etkin [56], the initial position of the CG was estimated by calculating the weight and position of each component within the aircraft. The component weights of electric aircraft can be considered constant, which allows for a more accurate estimate of the initial CG position compared to traditional fuel-powered aircraft. Figure 13 shows the internal layout of the AVM and the respective locations of the center of gravity for each component. The weight (W) and generated moment (M) of the primary aerodynamic surfaces, along with the flight components in the X and Y axes, are summarized in

Table 4. Weight and moment of the AVM components.

Group	W [N]	CG in X [m]	CG in Y [m]	M in X [N·m]	M in Y [N·m]
Gravity Center	1962.4	1.721	0.356	3377.29	698.614
Wing	262	1	0.3	262	78.6
Empennage	79.5	4.84	0.29	384.78	23.055
Tilt-rotor mechanism 1	314	0.97	0.31	304.58	97.34
Tilt-rotor mechanism 2	314	4.84	1	1519.76	314
Camera	48.1	0.53	0.19	25.493	9.139
Fuselage	479.7	1.25	0.31	599.625	148.707
Parachute	51.2	1.5	0.38	76.8	19.456
Avionics	11.7	1.09	0.28	12.753	3.276
Mass spectrometer	49	0.12	0.25	5.88	12.25
Batteries	353.2	1.07	0.25	377.924	88.3

.

The center of gravity of the AVM is calculated using Equation (6):

$${\left(X,Y\right)}_{CG}={\displaystyle \frac{\sum W·d_{(X,Y)}}{\sum W}}$$

(6)

where $d$ represents the distance of the $X$ component from the reference plane (the aircraft nose).

The location of the focal point determines the longitudinal static stability of the aircraft. If the center of gravity is located ahead of the focal point, the aircraft will be statically stable. The aerodynamic center of the wing or horizontal tail (typically located at 25% of the mean aerodynamic chord of the wing) is calculated using Equation (7):

$$X_{ac}=X_{LE}+0.25·MAC$$

(7)

where $X_{LE}$ is the reference point of the leading edge of the wing or tailplane (with the fuselage nose as the origin), and $MAC$ is the mean aerodynamic chord of the wing or horizontal tail.

The overall focal point of the aircraft is the weighted average position of the lift increments from all components. This requires consideration of the lift contributions from both the wing and tailplane, as well as the downwash effect [51], and can be calculated using Equation (8):

$$X_{focus}={\displaystyle \frac{C_{L\alpha ,wing}·X_{AC,wing}+C_{La,tial}·{\eta }_{tail}·\left(1-\frac{\partial \epsilon }{\partial \alpha }\right)·X_{ac,tail}}{C_{L\alpha ,wing}+C_{La,tial}·{\eta }_{tail}·\left(1-\frac{\partial \epsilon }{\partial \alpha }\right)}}$$

(8)

where $C_{L\alpha ,wing}$ are the lift contributions from the wing and horizontal tail, $X_{ac,wing}$ and $X_{ac,tail}$ are their aerodynamic center positions, ${\eta }_{tail}$ is the ratio of dynamic pressures at the tail and wing, and ${\displaystyle \frac{\partial \epsilon }{\partial \alpha }}$ is the derivative of the downwash angle with respect to the angle of attack. The calculated result shows that the AVM ‘s center of gravity is less than the specified value, indicating that the aircraft possesses good stability and handling qualities.

3.7. Performance Analysis

To ensure that the mission requirements for low-altitude flight operations are met by the propulsion system, an analytical investigation of flight performance metrics is conducted in this section. The system’s performance is evaluated across various flight conditions, and its impact on key factors such as stability, efficiency, and overall mission success is examined. Additionally, critical flight performance parameters are analyzed to determine whether the desired outcomes can be achieved by the propulsion system during low-altitude flight operations.

3.7.1. Velocity

Under steady-speed rectilinear flight conditions, the force equilibrium equations governing the AVM’s static aerodynamic balance are mathematically derived as follows:

$$T=D={\displaystyle \frac{1}{2}}\rho V^{2}S{C}_D$$

(9)

$$W=L={\displaystyle \frac{1}{2}}\rho V^{2}S{C}_L$$

(10)

where $T$ is static thrust, $D$ is aerodynamic drag, $L$ is lift force, $W$ is aircraft weight.

By solving these coupled equations simultaneously, the required thrust ($T_R$) can be derived as Equation (11):

$$T_R={\displaystyle \frac{W}{C_L/C_D}}$$

(11)

Figure 14 illustrates the variation in thrust requirements for the AVM at different flight speeds, along with the changes in propulsion system efficiency as thrust varies. The design flight speed of the AVM is constrained by the available thrust range and efficiency characteristics of the propulsion system. Through systematic evaluation of these constraints, the optimal cruise velocity and maximum attainable velocity are quantitatively determined [57,58]. As power increases, the flight speed exhibits a significant linear growth until it stabilizes at a certain power level (around 100 kW). This indicates that, in the high-power range, the increase in flight speed is limited by factors such as aerodynamic drag and propulsion system efficiency, making further linear growth difficult. Flight efficiency shows an inverse relationship with power, particularly at lower power levels, where efficiency increases more noticeably. At approximately 27 kW, efficiency reaches its maximum value, suggesting that the propulsion system operates most efficiently at this power level, and the AVM is in its optimal flight speed range (around 216 km/h, in the subsequent computational analysis, taking into account the drag effects of factors such as the propeller and passenger cabin, a cruise speed of 150 km/h is recommended.). Although higher power inputs can increase flight speed, the improvement in efficiency becomes limited, and excessive power may lead to energy wastage.

3.7.2. Range and Endurance

This section investigates the maximum range and flight time achievable by the aircraft with a given battery energy capacity. For distributed electric-driven tiltrotor aircraft, due to their characteristics of low-altitude flight and vertical takeoff and landing, the cruise condition range is defined as the $Range$, and the cruise time is defined as the endurance time $t_{cruise}$. The corresponding calculation equations are as follows [59]:

$$Range={\displaystyle \frac{E_{bat1}·{\eta }_p^2}{\frac{1}{2}\rho ·V_{cruise}^2·S·C_D}}$$

(12)

$$t_{cruise}={\displaystyle \frac{E_{bat1}·{\eta }_p}{\frac{1}{2}\rho ·V_{cruise}^3·S·C_D}}$$

(13)

where $E_{bat1}$ is the maximum energy from the battery system used for cruising, ${\eta }_p$ is the efficiency of the propellers, $\rho$ is the air density at the cruising altitude, $V_{cruise}$ is the cruising velocity, $S$ is the wing area, $C_D$ is the drag coefficient.

3.7.3. VTOL State

The distinctive feature of tiltrotor aircraft lies in their ability to perform both the flight functions of a conventional fixed-wing aircraft and the vertical takeoff and landing capabilities. This section analyzes the VTOL performance of the AVM by examining its maximum hover time and vertical climb rate during takeoff.

The maximum hover time of the AVM is given by Equation (14):

$$t_{hover}={\displaystyle \frac{E_{bat2}}{\frac{n·T^{3/2}}{\sqrt{2\rho \pi R^2}}·\frac{1}{{\eta }_{sys}}}}$$

(14)

where $E_{bat2}$ is the maximum battery energy available for vertical takeoff and landing, $n$ is the number of propellers, $T$ is the thrust produced by a single propeller, $R$ is the radius of the propeller, ${\eta }_{sys}$ is the overall efficiency of the propeller power system.

The maximum vertical climb rate of the AVM is determined based on the maximum available power of the propulsion system, as given by Equation (15):

$$V_c={\displaystyle \frac{\frac{n·T^{3/2}}{\sqrt{2\rho \pi R^2}}·\frac{1}{{\eta }_{sys}}}{W}}$$

(15)

where $P_{avail}$ is the maximum available power of the propulsion system, $W$ is the weight of the aircraft.

3.8. Design Results

The final design yields a heavy-lift tiltrotor unmanned aerial vehicle module integrated with a fixed-wing configuration. The key performance specifications are summarized as follows:

•

Total endurance time: greater than 40 min

•

Maximum takeoff weight: 300 kg

•

Maximum hover time: 5 min

•

Maximum vertical climb rate: 25 m/s

•

Service ceiling: 1000 m

•

Maximum cruise range: 80 km

•

Stall speed: 17.5 m/s

•

Cruise speed: 150 m/s

Three views and an isometric view of the prototype of the AVM are shown in Figure 15.

4. Numerical Simulation Method

In the previous chapters, the aerodynamic characteristics of the AVM were analyzed under design constraints. The numerical simulation studies conducted for the AVM are presented in this chapter. The results obtained from the simulations are analyzed to assess the effectiveness of the proposed design and to provide a basis for further refinement.

4.1. Numerical Calculation and Validation

Due to convergence difficulties associated with steady-state simulations of complex aerodynamic flows, ANSYS Fluent (Version 2020R1, a licensed software purchased by University of Science and Technology Beijing.) was used to solve the Reynolds-Averaged Navier–Stokes (RANS) equations. The Spalart–Allmaras (S-A) turbulence model was selected, given its relatively low sensitivity to mesh quality, superior numerical robustness near complex geometries, and reliable performance in predicting low-speed flow phenomena such as trailing-edge separation and ground effect-induced pressure gradients. The applicability of the S-A model has been extensively validated in the aerospace field through both numerical simulations and experimental data. Several studies, such as those by Tong, Diskin, and Tarsia Morisco, have confirmed the reliability of the model under various flow configurations [60,61,62]. These studies demonstrate that the S-A model excels under different flight conditions, providing a solid experimental foundation for its application in aerospace engineering [63]. Thus, it was selected for simulating the aerodynamic performance of the AVM across a range of angles of attack at its cruise speed.

In CFD, mesh generation is a crucial and resource-intensive preprocessing step that significantly affects solution accuracy and stability. To accommodate the complex geometry of the UAV, an unstructured tetrahedral mesh was used, which is particularly effective for irregular shapes due to its flexibility in node arrangement. Mesh refinement was applied non-uniformly—regions requiring detailed flow resolution, such as wing-fuselage junctions, wing and tail surfaces, were locally refined using curvature-based controls. The far-field was discretized with coarser elements to reduce computational cost and memory usage. Grid orthogonality was carefully preserved to mitigate skewness and enhance simulation accuracy. Boundary layer refinement was applied using a “body of influence” (BOI) strategy to better capture near-wall flow behavior and pressure gradients across the computational domain.

The boundary conditions were configured as follows: a velocity inlet with a defined magnitude and direction (33.33 m/s), a pressure outlet, and a symmetry condition along the longitudinal midplane due to the half-model setup. The fluid was modeled as an ideal gas, and no-slip wall conditions were applied to all solid surfaces. A second-order upwind scheme was used for spatial discretization, and the flow field was solved using the SIMPLEC algorithm. The computational domain, as shown in Figure 16a, spans 7.42 m × 35 m × 13 m. The full aircraft model was discretized in both space and time with sufficient mesh density. The Navier–Stokes equations were solved under the assumption of incompressible and steady-state flow. After mesh quality enhancement, the maximum grid skewness was reduced to 0.79, meeting accepted accuracy thresholds. Once the surface mesh achieved the required quality, the volume mesh was generated and filled with unstructured tetrahedral cells. Figure 16b shows the final surface mesh and locally refined regions.

To validate the reliability of the CFD methodology, a benchmark case based on the AIAA Drag Prediction Workshop II configuration—the DLR-F6 wing–body model—was simulated. Experimental data for this configuration were obtained in the 1990s using the ONERA S2MA wind tunnel (1.77 m × 1.75 m) in France, providing a reliable dataset for verification purposes. In this study, the DLR-F6 model was simulated at an angle of attack of 0.49° under the reference conditions specified in [64,65]. The computed lift and drag coefficients (C_L = 0.4712, C_D = 0.0307) were compared with the experimental values (C_L = 0.49, C_D = 0.0292). The relative errors of 3.8% for lift and 4.9% for drag fall within the acceptable 5% range, thereby confirming the credibility of the adopted CFD approach for simulating steady-state aerodynamic characteristics at small angles of attack.

4.2. Grid Independence Tests

To ensure that the numerical results were independent of mesh density and to optimize computational cost, a mesh independence study was performed. The objective was to determine the minimum mesh resolution capable of producing accurate and stable results for the given aerodynamic model. Eight mesh configurations (MK1–MK8), ranging from 800,000 to 3,000,000 elements, were generated for the AVM model. An unstructured tetrahedral mesh was created for the entire domain. All simulations were carried out at 0° angle of attack under cruise conditions.

Table 5. Grid independence verification.

	Face	cell	CL	CD	Lift (N)	Drag (N)	L/D
MK-1	34,210	811,781	0.729455	0.063167	2803.301	242.7504	11.54808
MK-2	41,250	964,602	0.735437	0.061001	2826.292	234.4273	12.05615
MK-3	54,230	1,234,966	0.738283	0.059429	2837.226	228.3857	12.42296
MK-4	55,106	1,237,245	0.738051	0.059449	2836.337	228.4612	12.41496
MK-5	66,586	1,430,889	0.738417	0.059333	2837.745	228.0159	12.44538
MK-6	87,592	1,840,163	0.743138	0.057945	2855.886	222.6823	12.82494
MK-7	119,358	2,459,404	0.74093	0.058213	2847.401	223.7135	12.72789
MK-8	155,704	3,121,630	0.74321	0.05707	2856.164	219.3195	13.02284

summarizes the simulation results for each configuration, and Figure 17 illustrates the variation in lift (C_L) and drag (C_D) coefficients as a function of mesh size.

As shown in Figure 16, beyond 2.45 million cells, both C_L and C_D values stabilize, indicating that further mesh refinement yields negligible changes. Thus, a mesh with approximately 2.5 million cells is selected as the baseline configuration, balancing accuracy with computational efficiency.

5. Numerical Calculation Results

5.1. Aerodynamic Characteristics

This section presents an analysis of the aerodynamic characteristics of the AVM aircraft, focusing on its lift coefficient (C_L), drag coefficient (C_D), C_L/C_D ratio, and drag polar, to evaluate whether the aerodynamic performance meets the design requirements and objectives.

According to the results shown in Figure 18a, the CFD simulation predicted a stall angle of α = 13°. For low-altitude flight and aircraft without lift-enhancing features, this value is considered relatively high. It indicates that the aircraft was able to maintain relatively high lift and aerodynamic performance even when approaching stall, without immediately entering a stall condition. The flight control system was shown to be capable of recovering from this state through active control, thereby providing greater flexibility, especially in environments requiring rapid response (e.g., emergency turns or obstacle avoidance). In addition, a higher stall angle indicates improved stability during low-speed, high-angle-of-attack flight, allowing the aircraft to better adapt to varying aerodynamic conditions across different flight regimes. This helped prevent significant performance degradation due to unexpected stalls, which is particularly crucial for low-altitude missions operating in complex terrain and adverse weather conditions.

Figure 18b shows the drag coefficient results. The CFD analysis revealed a relatively high C_D compared to typical small civilian aircraft, in which the drag coefficient usually ranges from 0.02 to 0.05. This was attributed to the blunt aerodynamic design of the aircraft’s fuselage and battery compartment, which featured a flat, blunt nose rather than a sharp cone. This design typically induced strong shock waves, resulting in steep pressure gradients and temperature variations around the aircraft, as illustrated in Figure 19. However, this shape effectively dispersed airflow pressure, reduced thermal accumulation, and lowered the aircraft’s surface temperature. The C_D at zero angle of attack was found to be 0.0577, primarily due to the high curvature of the wing airfoil. Even at negative angles of attack, the aircraft was still able to generate sufficient lift via the auxiliary and tiltrotor propellers, thereby ensuring stable flight and preventing excessive lift that might hinder landing.

Figure 18c presents the relationship between aerodynamic efficiency and angle of attack. The maximum lift-to-drag ratio occurs at α = 2°, where C_L/C_D = 13.4. This result indicates that a 2° angle of attack represents the most stable and energy-efficient configuration for the aircraft. At this point, lift and drag are relatively balanced, resulting in the highest aerodynamic efficiency, which is optimal for steady cruising and contributes to extended range. For most aircraft, the best lift-to-drag ratio typically occurs near 0° angle of attack. Between 0° and 5°, the aerodynamic efficiency of the AVM exhibited only slight variation—it first increased and then decreased. The lift-to-drag ratio peaked at 2° with a value of 13.4, and then dropped to 12.72 at 5°, indicating a relatively modest fluctuation. This suggests that at small angles of attack, the aircraft maintains good aerodynamic efficiency, and that the minor efficiency loss may be attributed to the increase in lift coefficient, which also leads to increased drag. At a 5° angle of attack, the aircraft experienced a significant increase in lift while still maintaining adequate aerodynamic efficiency, making it suitable for short-range vertical climbs during low-altitude operations. However, as the angle of attack increased further, aerodynamic efficiency declined sharply, primarily due to flow separation, increased induced drag, and amplified pressure differences, which together resulted in a marked increase in drag. This behavior is particularly critical for low-altitude aircraft, as larger angles of attack limit climb performance and maneuverability. Therefore, selecting a 2° angle of attack as the optimal point for aerodynamic efficiency helps optimize the cruise phase, reduces energy consumption, and extends endurance.

Figure 18d shows the drag polar for the entire AVM design process. Low-altitude aircraft operate in environments characterized by higher air density and complex airflow patterns, requiring strong aerodynamic performance to ensure stable and efficient flight. In the low-lift coefficient range (C_L < 1), the aircraft exhibited a lower drag coefficient, meaning it was able to maintain high flight efficiency with relatively low energy consumption—especially during cruise. This allowed the aircraft to sustain steady flight at low power, effectively extending endurance and improving overall efficiency. In these low-lift conditions, the aircraft demonstrates minimal drag, enabling stable operation at small or low angles of attack—particularly under turbulent or rapidly changing airflow conditions. Lower drag coefficients help mitigate the effects of aerodynamic instability, which is critical for low-altitude stability. During cruise, the lift coefficient remains in the low-lift region, keeping drag relatively low and thereby ensuring good aerodynamic performance and enhancing cruise efficiency.

According to Figure 18d, when the lift coefficient exceeds 1, the drag coefficient increases rapidly, particularly near C_L = 2, indicating the onset of stall. This result suggests that, to avoid stall during low-altitude flight, the AVM design should maintain a cruise lift coefficient near 1, thus avoiding high-lift regions and minimizing drag to preserve flight stability. Based on the drag polar trend, the aircraft should ideally operate within the low-lift region during cruise.

Finally,

Table 6. Comparison of aerodynamic coefficient.

Aerodynamic Parameters	Numerical
Maximum CL/CD	13.40
α for (CL/CD) MAX	2°
CL (α = 0°)	0.7411
CD (α = 0°)	0.0577
Maximum CL	2.0386
α for CLMAX	13°

summarizes the key results of the entire aerodynamic design process.

5.2. Aerodynamic Effects

This section systematically analyzes the aerodynamic effects of the low-altitude aircraft at various angles of attack by presenting the “static pressure,” “velocity magnitude,” “wall shear stress,” and “Vortex Core Region.” Through numerical simulation results, the effects of different angles of attack on the surface pressure and friction distribution, as well as flow separation and reattachment, are visually observed. Figure 19, Figure 20, Figure 21 and Figure 22 show the airflow velocity streamlines, static pressure contours, vortex core region contours, and surface frictional force contours around the AVM, with the angles of attack of the AVM being α = 0° (Figure 19), α = 8° (Figure 20), α = 15° (Figure 21), and α = 20° (Figure 22).

The static pressure distribution shown in Figure 19a reveals a relatively uniform static pressure profile across the AVM surface. High-pressure regions are observed at the nose and wing root, while the low-pressure regions are located at the trailing edge of the wings and the rear fuselage. The wide positively pressurized area near the nose contributes significantly to lift generation. The low-pressure area towards the rear corresponds to airflow acceleration and changes in streamline curvature, which are typical of stable flow over the AVM’s surface. The velocity magnitude streamlines confirm the smooth flow over the AVM, especially along the wing surfaces. There are no signs of significant flow separation or turbulence, indicating that the airflow remains stable and well-controlled. The converging flow near the trailing edge of the wings further indicates the absence of major flow disruptions. The smooth and steady flow behavior across the aircraft reduces drag, which is favorable for low-speed cruising and low energy consumption.

The Vortex Core Region contour map provides valuable insight into the rotational flow characteristics around the AVM. Figure 19b shows relatively low levels of vortex intensity at the tail and wing sections, which suggests that vortex shedding and strong turbulence are minimal. The absence of prominent vortex structures in this region further supports the conclusion of stable flow behavior, aligning with the smooth streamlines observed in Figure 19a. The minimal vortex formation also implies low energy loss due to turbulence, contributing to overall aerodynamic efficiency during cruise conditions.

The wall shear stress contour map highlights the frictional forces between the AVM surface and the airflow. The areas of higher wall shear stress are concentrated at the nose and wing root, where the airflow is accelerating, as expected in regions of high pressure. However, the shear stress values are relatively moderate across the rest of the aircraft surface, indicating efficient flow control and low drag. The absence of high shear stress values over the wing surface suggests that the AVM’s design successfully prevents excessive frictional losses, which would otherwise contribute to higher induced drag.

The static pressure contours show a slight increase in pressure difference between the upper and lower surfaces of the wing as the angle of attack increases to 8°. This indicates a mild increase in lift generation compared to the previous angle of attack (α = 0°). The pressure distribution around the fuselage and wing junction remains smooth, suggesting that the flow is still attached without significant pressure losses or flow detachment. The streamlines reveal a slight increase in curvature compared to the 0° configuration. However, the flow remains largely attached to the surface of the wing and fuselage, demonstrating stable airflow over the aircraft surfaces. The slight divergence near the leading edge of the wing indicates a mild adverse pressure gradient, but it does not lead to flow separation. Figure 20a suggests that the AVM is still in a stable flight regime, capable of maintaining smooth airflow necessary for stable lift production at low speeds.

Figure 20b indicates the presence of minor vortex activity, primarily around the wing and tail sections. The absence of strong vortex core formation implies that there are no significant flow separations or turbulent wake structures. This further supports the stability of the flow around the AVM at this angle of attack. The relatively low vortex intensity is indicative of controlled, smooth flow, which is favorable for maintaining stability and minimizing drag. The well-controlled vortex activity also reduces the chances of sudden disturbances such as pitching or yawing moments, further ensuring the aircraft’s stability during low-speed flight.

Figure 20c reflects the frictional forces acting on the aircraft surface due to the airflow. At α = 8°, there is a moderate increase in wall shear stress compared to the 0° configuration, particularly near the wing root and fuselage junction, where the flow accelerates. This indicates that the adverse pressure gradient is beginning to have an effect on the frictional forces. However, the shear stress values are still within reasonable limits, suggesting that drag is still under control, and the aircraft maintains efficient aerodynamic performance. The smooth distribution of shear stress on the wing surface suggests that there is no significant flow separation or excessive turbulence that would lead to drag spikes.

The static pressure contours shown in Figure 21a reveal a significant pressure differential between the wing and tail surfaces. The pressure near the leading edge of the wing remains relatively high, while the pressure in the aft region, especially at the wing trailing edge and rear fuselage, decreases significantly. This results in a localized region of negative pressure. The pressure difference between the upper and lower surfaces of the wing enhances lift generation, but the negative pressure at the rear signifies the onset of flow separation. The velocity streamlines show relatively high airflow speeds over the wing and tail surfaces. However, in the separated flow regions (indicated within the red box), there is a noticeable decrease in flow velocity. The non-uniform velocity distribution suggests the presence of accelerated or decelerated areas, which contribute to an increase in induced drag and a reduction in lift. The turbulent flow observed in the red-boxed area further supports the presence of instability due to flow separations.

Figure 21b indicates that vortex activity is concentrated around the wing and fuselage junction, especially in the red-boxed area. The presence of strong vortex cores near the wing root and fuselage suggests significant flow detachment and the formation of vortices. These vortices result in turbulence and flow instability, contributing to a loss of lift and an increase in drag. The relatively high vortex intensity further emphasizes the flow instability, which is a characteristic of the stall region.

Figure 21c reflects the frictional forces on the AVM’s surface. High shear stress is concentrated near the leading edge of the wing and the fuselage junction, where the flow is accelerating. In the red-boxed region, the shear stress values are lower, indicating that the flow has started to separate, resulting in a reduction in frictional forces. The smoother shear stress distribution in the tailplane region suggests that the flow remains more stable, assisting in the prevention of flow separation on the tail and helping to maintain aerodynamic efficiency.

At α = 15°, the AVM demonstrates good aerodynamic performance despite entering a stall state. The tailplane plays a significant role in mitigating the effects of flow separation and ensuring the aircraft remains stable and controllable. The analysis confirms that the AVM’s design is effective in maintaining stability and providing stall delay, ensuring safe and efficient flight at high angles of attack.

The static pressure contours shown in Figure 22a reveal substantial pressure losses in the flow separation regions. The low-pressure zones near the wing leading edge, wing-fuselage junction, and wing-tail boom junction have grown significantly, reflecting the AVM’s struggle to maintain lift. The increased pressure difference between the upper and lower surfaces of the wing indicates that the flow is no longer attached to the wing’s upper surface. This results in a dramatic loss of lift and a sharp increase in drag. The pressure losses confirm that the aircraft is in a stalled state, and the aerodynamic efficiency is severely compromised. The streamlines clearly show the separation of flow from the wing surface, particularly in the red-boxed regions. The flow is no longer attached to the upper surface of the wing, and significant turbulence is observed downstream of the separation points. This indicates the onset of instability, with the flow being turbulent and leading to increased drag. The increase in drag is due to the large areas of separated flow, which increases the overall aerodynamic resistance of the aircraft. Although the tail section begins to experience some turbulence, the separation effect is less pronounced compared to the wing, allowing the tailplane to maintain some control authority.

Figure 22b shows a significant increase in vortex intensity around the wing-fuselage junction and the wing-tail boom junction, where the flow separation occurs. The strong vortices in these areas are associated with turbulent flow and the formation of wake vortices, which contribute to a substantial increase in drag. The vortex formation is a clear indication of flow instability, which is common during stall conditions. The localized vortex activity suggests that while flow separation is significant, it is still confined to certain regions, leaving the control surfaces like the ailerons and ruddervators operational.

Figure 22c indicates that the AVM’s surface experiences significant frictional forces due to the flow separation. In the regions where the flow separates (highlighted in the red box), the shear stress is relatively lower, suggesting reduced friction as the flow is no longer attached. However, the shear stress remains relatively high at the leading edge and fuselage junctions where the flow is still attached, but it is clearly affected by the developing separation. The shear stress in the tailplane region remains relatively smooth, suggesting that the tailplane’s flow remains more stable compared to the wing, aiding in the aircraft’s stability.

5.3. Summary of Numerical Calculation Results

1. 

High Stability and Adaptability in Low-Altitude Flight

According to the simulation results, the AVM maintained high lift and low induced drag during low-altitude flight, especially at angles of attack ranging from 0° to 8°. This design ensured that the AVM exhibited excellent stability in low-speed, high-angle-of-attack flight conditions, effectively preventing stall caused by flow separation. This stability was particularly advantageous in complex terrains and adverse weather conditions, providing high safety and reliability for low-altitude missions such as urban air mobility.

2. 

Outstanding Aerodynamic Efficiency in Low-Altitude Flight

CFD simulations showed that the optimal lift-to-drag ratio occurred at an angle of attack of 2°, with a C_L/C_D value of 13.4, indicating the most efficient aerodynamic configuration for the AVM. Although the AVM maintained high lift and low drag at smaller angles of attack, an increase in angle beyond 5° led to higher lift but also increased drag. By optimizing the angle-of-attack range, the AVM could maintain lower energy consumption during the cruise phase, thereby enhancing endurance and meeting the design requirements for long-duration low-altitude flight.

3. 

Higher Stall Angle Ensures Stability During Low-Speed, High-Angle Flight

The stall angle was predicted to be 13°, which was relatively high compared to typical small civilian aircraft. This meant the AVM could maintain good aerodynamic performance even in high-angle-of-attack flight, avoiding immediate stall. Especially in emergency maneuvers or obstacle avoidance, the higher stall angle provided greater operational flexibility, ensuring stable flight even under high-angle-of-attack conditions. This characteristic was crucial for low-altitude flight, where rapid and stable responses were necessary.

4. 

Tailplane Design Effectively Delays Stall, Enhancing Aircraft Control

At an angle of attack of 15°, the AVM’s main wing experienced significant flow separation, leading to a decrease in lift and an increase in drag. However, the tailplane design effectively delayed flow separation and, by guiding the airflow, reduced the risk of stall. In low-altitude flight, the tailplane’s stability protection helped the AVM maintain good controllability, allowing rapid recovery from near-stall conditions and preventing large attitude changes. This design improved stability and maneuverability, ensuring the aircraft met the rapid response and high safety requirements for urban air mobility.

6. Discussion

The aerodynamic analysis indicates that the AVM generates sufficient lift with minimal drag, reduces energy losses caused by flow separation and turbulence, and thereby optimizes overall aerodynamic efficiency, which is crucial for maximizing endurance and stability during low-speed cruising. The AVM’s control surfaces, such as the wings, horizontal tailplane, and auxiliary rotors, remain effective even under low-speed and near-stall conditions. The smooth flow over the wings and the absence of strong vortex formations further confirm that the control surfaces provide adequate stability and control throughout the entire flight envelope, with the ability to recover quickly from disturbances such as gusts or sudden airflow changes. This characteristic is especially beneficial for tasks requiring precise control during low-speed flight, such as navigating complex terrain or performing vertical takeoff and landing. The aerodynamic design of the AVM offers a robust foundation for low-altitude, low-speed operations, delivering advantages in energy efficiency, flight stability, and control. The combination of efficient lift generation, low induced drag, and smooth airflow characteristics ensures that the AVM meets its performance objectives while minimizing energy consumption and maximizing flight stability.

Some performance comparisons with other aircraft designed for analogous missions were conducted. Tao Zhang et al. [66] from the University of Glasgow’s School of Engineering proposed the Skybus, a heavy-lift six-rotor eVTOL aircraft with a maximum passenger capacity of 30, aiming to introduce the “park-and-ride” concept to urban air mobility. Compared to smaller eVTOL configurations, the Skybus offers a promising approach for future urban public air transportation. However, its large size introduces additional challenges in terms of vehicle performance, dynamic behavior, and acoustic characteristics. The presence of multiple wings on the same horizontal plane causes aerodynamic interference due to vortex generation at the leading-edge wingtips, thereby affecting lift and drag distribution across the airframe. To mitigate these effects, the front and rear wings must be designed with specific anhedral and dihedral angles, which increase structural complexity and manufacturing costs. Moreover, the large passenger load poses considerable challenges for the energy system, and the feasibility of employing emerging technologies such as pure electric or fuel cell systems remains under discussion.

Mohsen Rostami et al. [67] from the Department of Aerospace Engineering at Toronto Metropolitan University developed a multidisciplinary possibility analysis approach to optimize the design of an eVTOL tilt-wing aircraft. Their method integrated multiple disciplines, including aerodynamics, propulsion, mass distribution, and stability and control, using MAPLA software (v2.1.7). While this method proved effective in concept generation, the resulting aircraft model was relatively coarse, with the aerodynamic performance of critical components such as propellers and wings not visually evaluated.

In contrast, the AVM proposed in this study provides a more compact and experimentally verifiable design solution. Its configuration—featuring tiltrotors combined with a high-aspect-ratio fixed wing—offers improved stall resistance and enhanced aerodynamic stability in low-altitude flight conditions. Furthermore, the 100 kg class design simplifies manufacturing, testing, and iterative refinement, thereby enabling practical validation of advanced aerodynamic and control theories in realistic urban and mountainous environments.

Therefore, the AVM, as a prospective urban air mobility platform, requires further refinement to optimize its aerodynamic performance. Based on the results presented in this paper, the AVM demonstrates excellent aerodynamic characteristics and mission adaptability, and remains a promising candidate for future UAM applications. Figure 23 shows an artistic rendering of the AVM operating in a low-altitude environment.

7. Conclusions and Future Works

The conceptual design and aerodynamic analysis of a tiltrotor aerial vehicle module (AVM) used in split-type flying cars have been presented in this study. The design process incorporated multi-constraint optimization based on flight missions, payload, and energy efficiency. The AVM concept was validated using CFD simulations, showing strong alignment with performance targets.

Key outcomes of this work include:

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High aerodynamic efficiency: The AVM achieves a maximum lift-to-drag ratio of 13.4 at 2° angle of attack, ensuring optimal cruise performance.

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Robust low-speed stability: The stall angle reaches 13°, exceeding that of typical small UAVs, enabling safe operation in urban and mountainous low-altitude environments.

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Vertical takeoff and hover capabilities: The AVM supports a maximum hover time of 5 min and a vertical climb rate of 25 m/s, meeting VTOL mission requirements.

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Compact and practical design: With a maximum takeoff weight of 300 kg, the AVM strikes a balance between payload capacity and manufacturability, making it suitable for experimental validation.

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Enhanced flow control and stall resistance: Tailplane and rotor placement are optimized to delay stall onset and maintain control authority in near-stall conditions.

While the AVM shows promising aerodynamic performance based on theoretical calculations and CFD simulations, further refinement is necessary. Future work will focus on experimental validation through scaled wind tunnel testing and the incorporation of more complex urban flight conditions, such as heat island effects and surface gusts, to assess unsteady flow interactions. A detailed investigation of the electric propulsion system—including thermal management, structural integration, and airworthiness compliance—will be essential, with consideration also given to hybrid-electric alternatives. Additionally, structural design considerations will play a crucial role in improving AVM performance. Specifically, efforts will be directed towards designing the wing and battery bay truss structures, as well as performing topology optimization on specific components to achieve weight reduction and enhanced performance. The use of materials such as magnesium alloys and carbon fiber composites for critical structural components, including high-speed electric motors and airframe skin, will further optimize the overall AVM performance. Moreover, as this study remains within the conceptual design stage, future efforts will expand into structural analysis, control surface sizing, and dynamic stability evaluation, providing a more comprehensive basis for real-world implementation.