In this study, two representative eVTOL configurations developed by NASA—the lift + cruise [8] and the tiltrotor [11]—are selected as the reference aircrafts for comparative analysis. Geometry sizing is conducted while maintaining target wing loading and disk loading, reflecting aircraft performance during both the cruise and VTOL phases. These parameters govern geometry adjustments throughout the sizing iterations, where the wing area is adjusted to sustain the target wing loading, while the aspect ratio, taper ratio, and trailing edge angle are kept constant. Tail size is adjusted using a fixed tail volume coefficient, based on the main wing area and fuselage length. Rotor size is determined by disk loading and MTOW, calculating the total required disk area, which is then used to determine the rotor radius.

The aerodynamic analysis is based on the methods proposed in [20]. Vortex lattice method (VLM) is employed to compute the wing lift. Parasite drag is evaluated through a component buildup approach, where the viscous drag of individual components is calculated using skin friction based models with form factor corrections and then summed in an area-weighted manner. Induced drag is derived from the lift distribution predicted by the vortex lattice solution, with an additional lift-dependent viscous correction. The total drag is constructed by combining parasite and induced contributions within a unified aerodynamic buildup framework [20]. The aerodynamic methods have been validated in the literature for conceptual aircraft analysis [21].

Figure 2 illustrates the architecture of the all-electric propulsion system, which is powered by batteries that supply electrical energy to the aircraft’s motors. The motors drive the rotors, with a battery capacity constraint ensuring that at least 20% of the battery’s capacity remains after completing a typical mission. The performance parameters for the electric propulsion system components are sourced from the literature [22–24]. While the efficiency of the propulsion system varies with operating conditions, a representative efficiency value is chosen for this study, reflecting typical mission conditions. The rotor performance such as thrust coefficient, power coefficient and efficiency is solved using XROTOR [25], a rapid solver for rotor aerodynamics based on blade element-vortex theory, which enables quick and reasonably accurate calculation of rotor performance.

The aircraft mass is estimated using a physics-based weight buildup method [20], adapted from the Airbus A³ Vahana conceptual study [26]. The approach is load-driven rather than regression-based, making it suitable for novel eVTOL configurations without historical statistical data. Structural components, including the wing, fuselage, and rotors, are sized according to limit load conditions and material strength constraints under prescribed safety factors. Propulsion and electrical system masses are determined based on maximum thrust requirements and transmitted power levels.

The mission profile is divided into two distinct phases: VTOL mode (hover, vertical climb, transition, and vertical descent) and fixed-wing flight mode (accelerated climb, and cruise). A simplified model estimates the required power for each flight mode. Hover power [27] is calculated as Equation 1: P h = T 3 / 2 2 ρ n A d · η e l e c · FOM (1) where T represents the thrust, ρ is the air density, n is the number of rotors, FOM is rotor figure of merit calculated using XROTOR and A d is the disk area. For the transition phase, power consumption is assumed to remain constant, averaging the starting and ending power values (Equation 2) [27]: P t r = P h + P h / κ 2 (2) where P h is the hover power, P t r is the transition power and κ = 10 is the power reduction factor.

In fixed-wing flight mode, the power is estimated using the lift-to-drag ratio ( L / D ) (Equation 3): P c r u i s e = W t o · g · V c r u i s e L / D · η e l e c · η p r o p (3) where W t o is total mass of the vehicle g is gravitational acceleration V c r u i s e is cruise velocity η e l e c is efficiency of the electric powertrain and η p r o p is propeller efficiency calculated using XROTOR.

The study uses the Fourth-Order Runge-Kutta method (RK4) for numerical integration, iterating through the following steps to solve the system at each time step h (Equation 4): y n + 1 = y n + h 6 k 1 + 2 k 2 + 2 k 3 + k 4 (4) where y n + 1 is new value of the state variable at the next time step y n is current value of the state variable, h is time step k 1 , k 2 , k 3 , k 4 is intermediate calculations for the Runge-Kutta method and the k is the function evaluation at each step. The RK4 method is applied to time-varying state variables, such as the aircraft’s altitude and energy consumption.

Rotors are the dominant noise sources of eVTOL aircraft. This study accounts for harmonic and broadband noise contributions. Harmonic noise is primarily generated by the periodic motion of the blades and consists of thickness noise and loading noise. The prediction is based on the rotor aeroacoustic formulations developed by Hanson [28] and Farassat [29]. The sound pressure fluctuations associated with the m-th harmonic order are calculated using Bessel functions and azimuth angle corrections. The total harmonic acoustic pressure SPL harmonic is integrated over the blade span, yielding the total SPL according to Equation 5: SPL harmonic = 20 ⁡ log 10 p harmonic p ref (5) where p harmonic is harmonic sound pressure and p ref is reference sound pressure.

Broadband noise mainly arises from the interaction between the turbulent boundary layer and the blade trailing edge. In this study, the trailing-edge noise model proposed by Li et al. [30] is adopted. The total sound pressure level is obtained by energetic superposition of the harmonic and broadband components (Equation 6): SPL total = 10 ⁡ log 10 10 SPL harmonic / 10 + 10 SPL broadband / 10 (6) where SPL b r o a d b a n d is broadband sound pressure level and SPL total is total sound pressure level.

The implementation details and validation basis of the present aeroacoustic framework follow Clarke [21], in which the harmonic, broadband, and isolated rotor noise predictions were validated against published experimental data. Finally, an A-weighting filter is applied to account for human auditory frequency sensitivity. In this study, the acoustic observer is located at a distance of 36 ft from the rotor axis, which is adopted as a near-field reference location commonly used in rotor aeroacoustic studies [9]. In addition, the rotor tip mach number is limited to 0.65 to avoid excessive compressibility effects and to suppress high-speed impulsive noise.

The aircraft sizing process begins with defining the mission profile and associated requirements. The sizing mission profile, based on NASA’s research on UAM missions [31], is illustrated in Figure 3.

The target payload and cruise speed are specified as 200 kg and 234 km·h^-1, respectively. The takeoff altitude is assumed to be 1800 m above mean sea level, while the cruise altitude is set to 1,200 m above ground level. During VTOL operations, the aircraft operates under nominal conditions with a vertical climb rate of 2.5 m·s^-1. In the fixed-wing flight regime, the climb segment is modeled using a constant climb flight-path angle of 5°. A detailed summary of the mission requirements is provided in Table 1.

Based on the results illustrated in Table 2 and Figures 4–8, a systematic configurational trade study is conducted to compare the lift + cruise and tiltrotor eVTOL configurations under identical UAM mission requirements. This comparison focuses on differences in aircraft weight distribution, energy efficiency, key design parameters, and acoustic performance.

Quantitatively, the overall weight breakdown of the two configurations shows broadly similar proportions across major subsystems, as illustrated in Figure 4. However, the tiltrotor achieves a lower MTOW (964 kg) compared to the lift + cruise (1,013 kg), resulting in a reduction of approximately 4.8%. This difference is primarily attributed to the fact that, in the lift + cruise, the vertical lift system—including lift rotors and supporting booms—becomes inactive dead weight during cruise, contributing additional parasite drag. The aerodynamic performance comparison under cruise conditions is shown in Figure 5, where the L/D trends indicate superior cruise efficiency for the tiltrotor configuration. The analysis was conducted at an altitude of 3,000 m, with a speed range of 0.1–0.3 Mach, and an angle of attack ranging from 0° to 10°. The lower cruise efficiency of the lift + cruise increases thrust and energy demand for the same mission.

In terms of mission energy consumption, the lift + cruise requires 75.7 kWh to complete the 200 km mission, whereas the tiltrotor requires 68.7 kWh, representing an energy reduction of approximately 9.3%. The higher energy requirement of the lift + cruise necessitates a larger battery mass, which in turn increases the MTOW, creating a positive feedback loop between energy requirement, aircraft weight, and thrust.

At the propulsion system level, both configurations exhibit distinct engineering trade-offs. The tiltrotor employs fewer but significantly larger rotors (3.29 m in diameter vs. 1.59 m for the lift + cruise), with a figure of merit (FOM) of 0.84, compared to 0.79 for lift + cruise. The rotor performance was analyzed at an altitude of 3,000 m and a cruise speed of 234 km·h^-1, with varying advance ratios under cruise conditions is presented in Figure 6. Under one-engine-inoperative (OEI) conditions, each tiltrotor must bear a higher load and provide greater redundancy, necessitating more robust blade and drivetrain designs. Consequently, the propulsion system accounts for a slightly higher fraction of MTOW (20% vs. 18%). In contrast, the lift + cruise achieves greater system redundancy through eight lift rotors, reducing the structural burden on individual rotors and enhancing fault tolerance.

The noise analysis reveals significant differences in acoustic performance between the two configurations. The acoustic directivity and harmonic–broadband composition are presented in Figure 7. The lift + cruise exhibits a sound pressure level (SPL) of 92 dB at a 45° observer location, notably higher than the 69 dB predicted for the tiltrotor. This disparity is primarily due to the high disk-loading rotors in the lift + cruise, which generate strong harmonic noise during VTOL operations, while broadband noise contributions remain relatively modest. Without dedicated acoustic optimization, the noise levels of the lift + cruise significantly exceed typical residential noise limits for urban air mobility (<67 dB) [32], posing a critical constraint for urban deployment.

Figure 8 presents the payload–range characteristics of the two configurations. Both curves exhibit an approximately linear trade-off, reflecting the energy-limited nature of fully electric propulsion, where increased range requires battery mass to replace payload under a fixed structural mass and usable energy constraint. At the 200 km design point, the two curves intersect. The slope of each curve is governed by the cruise energy consumption per unit distance. Owing to its higher aerodynamic efficiency, the tiltrotor requires less additional battery mass per unit range, resulting in a shallower slope and 2%–3% greater maximum range. For shorter ranges (<200 km), the tiltrotor shows slightly lower payload capacity due to its higher propulsion mass fraction and smaller baseline battery mass. Overall, the lift + cruise is more suitable for short-range missions, while the tiltrotor is advantageous for medium-to long-range operations.

Overall, the tiltrotor demonstrates superior performance in terms of aerodynamic efficiency, MTOW, mission energy consumption, and acoustic characteristics. However, its increased configurational complexity—particularly associated with rotor tilting mechanisms, mode transitions, and aeromechanical coupling—poses higher engineering challenges. Conversely, the lift + cruise offers a simpler structural and control architecture with higher inherent redundancy, but suffers from penalties in aerodynamic efficiency and noise performance. Therefore, the selection of an appropriate configuration must be guided by specific mission requirements and operational scenarios.

To evaluate the impact of battery technology advancement on the performance of different eVTOL configurations, a sensitivity study is conducted by varying the battery energy density from 300 Wh·kg^-1 to 700 Wh·kg^-1. As illustrated in Figures 9–11, increasing battery energy density leads to a pronounced reduction in MTOW, mission energy consumption, and characteristic power requirements for both configurations. The trends are particularly significant in the lower energy-density range (300–400 Wh·kg^-1), indicating a strong sensitivity of aircraft performance to battery energy density in this regime.

For the lift + cruise, the MTOW is reduced by approximately 37% over the considered energy-density range, while the tiltrotor exhibits a slightly larger reduction of approximately 40%. In both cases, the battery mass fraction decreases substantially with increasing energy density. Specifically, the battery mass fraction of the lift + cruise decreases from approximately 31%–15% of MTOW, while that of the tiltrotor decreases from 30% to 14%. This increase in battery energy density effectively reduces the battery mass required to complete the mission, thereby lowering the overall aircraft weight and further decreasing power requirements and mission energy consumption in a positive feedback loop.

Within the energy-density range of 300–700 Wh·kg^-1, the total mission energy consumption of the lift + cruise and tiltrotor decreases by approximately 32% and 36%, respectively. Higher battery energy density enables a substantial reduction in the required system power level, with the maximum power requirement decreasing by approximately 37% for the lift + cruise and 40% for the tiltrotor configuration.

Despite these absolute improvements, the differences in mission energy consumption, power and MTOW between the two configurations remain relatively constant across the entire energy-density range. This indicates that the inherent aerodynamic efficiency and system integration characteristics of each configuration are the primary factors governing overall performance, rather than the absolute level of battery technology.

In the higher energy-density regime (600–700 Wh·kg^-1), the rate of power reduction gradually diminishes. This trend suggests that once battery energy density is no longer the primary limiting factor, system power requirements become increasingly constrained by other factors, including the aerodynamic efficiency limits of the rotor and wing, propulsion system efficiency, and inherent disk loading and aerodynamic layout characteristics of each configuration. Consequently, the marginal performance gains achievable solely through further increases in battery energy density are expected to diminish, and future performance improvements will increasingly rely on configuration optimization, aerodynamic refinement, and enhancements in propulsion system efficiency.

Overall, the sensitivity analysis confirms that battery energy density is a crucial factor influencing eVTOL aircraft weight, mission energy consumption, and power requirements for both configurations. Although both configurations exhibit similar sensitivity to changes in battery energy density, the differences in MTOW, energy consumption, and power requirements attributable to configuration characteristics remain relatively stable across the energy-density range. At higher energy densities, the system’s performance shifts from being primarily energy-limited to being constrained by aerodynamic and configurational factors.