Helicopter climb/descent maneuvering control is mainly achieved by adjusting the rotor collective pitch in forward flight conditions. The aeroacoustic characteristics of helicopter rotors during aperiodic variations in collective pitch under forward flight conditions are calculated and analyzed based on the CFD method and FW-H equations. First, a set of analytical methods for aperiodic rotor aeroacoustic characteristics is established. The AH-1G rotor in forward flight and the NACA rotor under a ramp increase in collective pitch are simulated, and the numerical analysis method employed is validated by comparison with experimental data. Then, the aeroacoustic characteristics of the AH-1G rotor during a ramp increase in collective pitch are analyzed, with a focus on the sound pressure peak in the blade-vortex interaction state. Finally, a detailed analysis is conducted on the BO-105 rotor in forward flight during aperiodic variations in collective pitch. In addition, parameters such as the collective pitch change rate are quantified, leading to several conclusions: during ramp increases and decreases in collective pitch, positive and negative overshoot phenomena occur respectively, with both cases showing a significant increase in acoustic levels.

A Computational Fluid Dynamics (CFD) method for variable-diameter (VD) rotor is developed based on the unsteady Reynolds-averaged Navier–Stokes (URANS) equation and moving-embedded grid method to investigate the novel aerodynamic behaviors and the unsteady aerodynamic characteristics of rotor during diameter morphing process, such as spanwise motion of tip vortices, blade–vortex interaction (BVI), and vortex–vortex interaction (VVI). Comparative analysis is conducted to verify the unsteady aerodynamic analysis method in collective pitch ramp change state. The in-depth simulations of three-dimensional (3D) flowfiled during rotor diameter morphing process are carried out in hover and forward flight conditions, and the influences of diameter morphing amplitudes, speeds, and patterns on the unsteady aerodynamic characteristics are revealed. The results indicate that the dynamic change of rotor diameter can induce the aerodynamic inertial behaviors like “hysteresis effects” and “overshoot responses”, accompanied with flow coupling between the spanwise flow and the blade-tip vortices and leading to regional novel BVI and VVI phenomena in the midspan region of the blade. On the whole, the amplitude of diameter variation exhibits a certain linear relationship with the rotor thrust; the unsteadiness of the flow field, such as the overshoot of unsteady load, increases with the VD speed; the diameter change pattern can affect the peak and phase of the aerodynamic loads, and the change degrees of the aerodynamic load are different in the diameter increment and decrease cases. The conclusions can provide engineering guidance for the design of diameter morphing strategies of intelligent VD rotors.

Helicopters operating in high-speed forward flight conditions frequently generate High-Speed Impulsive (HSI) noise, posing significant challenges for effective noise control and limiting their operational flexibility. To address this critical issue, this study proposes a novel active noise reduction approach that dynamically adjusts rotor diameter, aiming to alleviate the generation and intensity of HSI noise. Utilizing the Chinese Laboratory of Rotorcraft Navier-Stokes (CLORNS) solver combined with the Ffowcs Williams-Hawkings (FW-H) acoustic analogy, aerodynamic and acoustic characteristics of the AH-1G helicopter rotor undergoing diameter adjustments were comprehensively simulated. This study methodically investigated key factors influencing the acoustic field, including the scale of rotor diameter retraction and the speed of retraction. The results clearly demonstrate the effectiveness of the proposed approach, achieving significant noise reductions ranging from 4 dB to 7 dB at crucial operational points while ensuring aerodynamic stability. Parametric analyses further reveal that the scale of rotor diameter retraction is a decisive factor in noise mitigation, and slower retraction speeds are beneficial, contributing to smoother flow transitions and diminished sound pressure fluctuations. Moreover, to overcome potential noise amplification induced by rotor control actions, this research introduces a compensation method incorporating a dynamic parameter adjustment mechanism. The developed control strategy is particularly effective in eliminating the noise growth observed in regions characterized by abrupt noise increases. Overall, this study provides an innovative and practical solution for helicopter noise management, substantially enhancing operational flexibility and acoustic comfort in high-speed flight conditions.

The rapid transition phase from the helicopter mode to the fixed-wing mode entails complex, highly unsteady aerodynamic phenomena induced by rotor tilting. Most prior studies adopt steady or quasi-steady assumptions, which fail to capture the transient aerodynamic characteristics throughout the continuous tilting process. To address these critical bottlenecks, this work investigates high-fidelity aerodynamic modeling and simulation for tiltrotor rapid transition maneuvers. A fully transient high-fidelity aerodynamic simulation framework is developed based on moving-embedded grids and computational fluid dynamics technology, and its validity is verified using the quasi-steady aerodynamic test data of the XV-15 tiltrotor and the transient aerodynamic test data of an experimental rotor. Leveraging the proposed framework, high-fidelity numerical simulations are performed to systematically analyze the evolution of aerodynamic characteristics and multi-scale vortex systems during the continuous dynamic tilting transition, and to examine the transient unsteady aerodynamic evolution of the rotor under coordinated control of tilt and collective pitch. The research results indicate that with the increase in the tilt angle, the rotor thrust decreases gradually, and the amplitude of transient thrust oscillation increases progressively. When the rotor tilts to the axial flow state, the periodic blade motion and high-intensity vortex clusters trigger self-excited oscillations, resulting in sustained and substantial oscillations of aerodynamic forces. Increasing the rotor collective pitch to compensate for the reduction in rotor effective angle of attack caused by tilting can alleviate the rotor thrust loss and effectively suppress the transient oscillation amplitude of rotor thrust simultaneously.

A novel TENO-adaptive accuracy (TENO-A) scheme is proposed to perform complex flowfield simulations containing abundant vortex field details and shock wave discontinuities. This hybrid scheme not only enhances the low dissipation property of the TENO scheme in smooth regions but also improves the discontinuity-resolving capability for shock waves while retaining the robustness of the scheme. A unified and robust discontinuous detector using the information within the global stencil width is applied to separate discontinuous and smooth regions, which is not case and parameter sensitive. In discontinuous regions, the THINC scheme with jump-like distribution is adopted to well approximate a discontinuity within a grid. Besides, in the smoothest region, the global reconstruction value is adopted as the final reconstruction value, and the accuracy is improved from the original fifth-order to sixth-order; in the general smooth region, based on the Kriging method, the candidate stencil reconstruction is employed for final reconstruction with nonlinear weights, which improves the accuracy of the candidate stencil from the original third-order to fourth-order. As a result, the dissipation is significantly reduced using both types of reconstruction, which is beneficial to resolve small-scale flow structures. A set of numerical results demonstrates that the TENO-A scheme performs better in one-dimensional and two-dimensional cases than the standard TENO scheme and is able to predict complex flowfield without the necessity of parameter tuning case by case, and the hybrid scheme can restore high-order accuracy, maintain low dissipation property, and avoid spurious oscillations.

In order to capture the formation and evolution of blade tip vortex and blade vortex interaction (BVI) events, the high-fidelity hybrid TENO8-THINC reconstruction scheme is proposed to simulate unsteady flowfield of rotor. The hybrid scheme ensures robustness and stability in capturing the discontinuities while minimizing numerical dissipation in the smooth regions dominated by the tip vortices, thus significantly improving the detail capture ability of vortex field and providing accurate aerodynamic input for noise prediction. The noise characteristics of BVI state and non-BVI state are newly investigated by decomposing Farassat 1A formula into thickness, load amplitude and load derivative terms, and the results demonstrate that the dominant factor of BVI noise is the load derivative term. Then, the sound pressure time histories of oblique and parallel interaction regions of OLS rotor in BVI state are calculated and compared, and it indicates that oblique interaction on the advancing side has the greatest influence on BVI noise. Finally, parameter analysis of the influence of anhedral tip blades on BVI noise is carried out, illustrating that anhedral blade divides the main blade tip vortex into two vortex systems with similar intensity, which alleviates the strength of oblique interaction on the advancing side and parallel interaction, so as to achieve the purpose of BVI noise reduction, and the reduction of sound pressure level is nearly 5 dB.

Based on the Ffowcs Williams–Hawkings equations and the computational fluid dynamics method, the rotor’s aeroacoustic characteristics, considering the influence of the downwash flowfield on the sound propagation process, are calculated and analyzed. First, a set of analysis methods for the aeroacoustic characteristics is developed, and a convection-based propagation time model is developed, where acoustic group velocity along source–observer lines quantifies flowfield effects. Then, the rotor’s aerodynamic and aeroacoustic characteristics are calculated, and the employed numerical analysis method is validated through the comparisons with experimental data. Finally, the aeroacoustic characteristics of the rotor in hover are analyzed, and the sound pressure positive peak point with/without the influence of the flowfield of downwash on the propagation time is discussed in detail. In addition, parameters, such as the rotor’s collective pitch and the azimuthal angle of the sound source, are quantified, and some conclusions are obtained. For those observers below the rotor rotation plane, the downwash flowfield will influence the sound propagation time, resulting in the increase in the sound pressure and the advance of the arrival time.

Generally, the performance advantages of rotor airfoils with high aerodynamic efficiency might be reduced when applied on rotors, however, the quantitative research on the three-dimensional effects of rotor that lead to this problem is still very lacking. In order to quantitatively investigate the influences of three-dimensional effects on the aerodynamic characteristics of rotor airfoils, a novel numerical strategy for analyzing equivalent angle of attack and aerodynamic shape of blade profiles is proposed based on high-precision CFD method and the inverse design method, and the simulations on equivalent profiles and their equivalent angles of attack are conducted at various tip Mach numbers and aspect ratios of rotor blades. The numerical results indicate that the equivalent profiles in hover are mainly affected by radial flow on rotor, and the difference of aerodynamic shapes between equivalent airfoils and blade profiles are larger with the section closer to blade tip. The major influence on equivalent angle of attack is the vertical induced velocity, and the loss of angle of attack at root is larger than at tip. Additionally, lower tip Mach number and larger aspect ratio of rotor blades help to reduce the variation of the equivalent airfoil and loss of angle of attack. The quantitative results could provide theoretical references for the design of high-performance rotor airfoils and their applications on rotors.

The control stability and accuracy of quad tiltrotor UAVs is improved when encountering external disturbances during automatic flight by an active disturbance rejection control (ADRC) parameter self-tuning control strategy based on a radial basis function (RBF) neural network. Firstly, a nonlinear flight dynamics model of the quad tiltrotor UAV is established based on the approach of component-based mechanistic modeling. Secondly, the effects of internal uncertainties and external disturbances on the model are eliminated, whilst the online adaptive parameter tuning problem for the nonlinear active disturbance rejection controller is addressed. The superior nonlinear function approximation capability of the RBF neural network is then utilized by taking both the control inputs computed by the controller and the system outputs of the quad tiltrotor model as neural network inputs to implement adaptive parameter adjustments for the Extended State Observer (ESO) component responsible for disturbance estimation and the Nonlinear State Error Feedback (NLSEF) control law of the active disturbance rejection controller. Finally, an adaptive attitude control system for the quad tiltrotor UAV is constructed, centered on the ADRC-RBF controller. Subsequently, the efficacy of the attitude control system is validated through simulation, encompassing a range of flight conditions. The simulation results demonstrate that the Integral of Absolute Error (IAE) of the pitch angle response controlled by the ADRC-RBF controller is reduced to 37.4° in comparison to the ADRC controller in the absence of external disturbance in the full-states mode state of the quad tiltrotor UAV, and the oscillation amplitude of the pitch angle response controlled by the ADRC-RBF controller is generally reduced by approximately 50% in comparison to the ADRC controller in the presence of external disturbance. In comparison with the conventional ADRC controller, the proposed ADRC-RBF controller demonstrates superior performance with regard to anti-disturbance capability, adaptability, and tracking accuracy.

The classical WENO schemes perform well for most flow field simulations, they may encounter the 'Cannikin Law' trap, that is, the lowest accuracy order of the scheme may have a significant influence on the simulation. In this article, a novel WENO scheme (termed HPWENO) with improved convergence order is proposed to alleviate this issue. The research in this article is structured around three key steps: Firstly, the stencil is classified as either smooth stencil or non-smooth stencil by using the classification strategy of the hybrid WENO scheme. Secondly, perturbed polynomial reconstruction with double free-parameters is proposed. Finally, the new reconstruction coefficients containing multiple free-parameters, built on the classical fifth-order WENO schemes, are obtained by using the perturbed polynomial reconstruction. Compared to the fifth-order WENO schemes, a maximum two-order of accuracy improvement in candidate stencils and one-order of accuracy improvement in global stencil can be achieved by adaptively adjusting the values of these free-parameters, resulting in sixth-order accuracy in global stencil and fifth-order accuracy in candidate stencils. Compared to the classical fifth-order WENO5-Z scheme and the WENO-AO(5,3) scheme, numerical examples show that the HPWENO schemes have higher convergence ratio, provide sharper solution profiles near discontinuities, and perform well in resolving small-scale structures. Compared to the sixth-order WENO-CU6 scheme and the seventh-order WENO7-Z scheme, the proposed HPWENO schemes outperform the two schemes in resolving the small-scale vortex of two-dimensional issues, and it saves approximately 15% and 25% of computational resources, respectively.

Influence of Mach numbers and telescoping positions (TP) on the aerodynamic-enhancement mechanism of Variable Diameter-and-Speed Tilt-Rotors (VDSTR) was investigated by utilizing Blade Element and Momentum Theory (BEMT), with the original metal-blade tilt-rotors of XV-15 employed as the benchmark. First, by fixing the operational tip Mach numbers while allowing free cooperation of variable diameter and speed, the mechanism was analyzed, and the effects of air compressibility and telescoping position on the mechanism were further scrutinized. Then, advantages over single-variable techniques were highlighted by comparing hover and high-speed forward flight performance. The results highlight distinct aerodynamic improvement mechanisms for VDSTR in different flight states. The reduction of induced power compensates for increased profile drag power, a fundamental mechanism enhancing VDSTR in hover. The mechanism relies on blade contraction in high-speed forward flight, effectively reducing profile drag power. Moreover, compared to the original XV-15 blade, VDSTR enhances forward flight efficiency by 6.93

In this paper, a fixed-time tracking control scheme based on the fixed-time disturbance observer is proposed for a 6-DOF unmanned aerial helicopter (UAH) with flight path constraints and composite disturbances. The composite disturbances are composed of external disturbances and system uncertainties. By using the hyperbolic tangent function, an improved disturbance observer is developed to estimate composite disturbances. Furthermore, a fixed-time back-stepping control method is employed for the position and attitude loops, which enables the UAH to track the expected path within the flight path constraints. Simulation results are given to demonstrate the effectiveness and advantages of the proposed scheme.

This paper describes the aerodynamic simulation and optimization of NACA 0012 airfoil at a low Reynolds number using unsteady Reynolds-averaged Navier-Stokes (URANS) and Spalart–Allmaras turbulence model in Ansys Fluent. The purpose of this paper is to simulate and optimize the airfoil to get better aerodynamic performances at low Reynolds numbers. The Parsec method was selected for the optimization of the NACA 0012 airfoil. Both of these airfoils are simulated using CFD Fluent between 0 to 13-degree angle of attack at a low Reynolds number of 200000. To simulate the airfoil, mesh generation is crucial so an O-grid structured mesh is created. After the simulation, several aerodynamic performances are compared between the airfoils, such as lift coefficient, drag coefficient, pressure coefficient, and lift-to-drag ratio. And the calculated results from Xfoil are taken as references. Between NACA 0012 and optimized NACA 0012, the optimized airfoil showed better aerodynamic performances than the normal one, which was the goal of this paper. Later on, the different flow field variables, such as density, temperature, pressure, and vorticity magnitude were analyzed and compared. Both the airfoils at a different angle of attack were analyzed for these functions, like 7°, 11°, and 20° AOA. During the analytical process, Q-criterion appears to be a very important method of vortex identification in the flow field. With this analysis, we came to know, that as the angle of attack increases the adverse pressure gradient also increases, which creates a big reverse flow.

To investigate the distinct properties of the helicopter rotors during circling flight, the aerodynamic and dynamic models for the main rotor are established considering the trim conditions and the flight parameters of helicopters. The free wake method is introduced to compute the unsteady aerodynamic loads of the rotor characterized by distortions of rotor wakes, and the modal superposition method is used to predict the overall structural loads of the rotor. The effectiveness of the aerodynamic and the structural methods is verified by comparison with the experimental results, whereby the influences of circling direction, radius, and velocity are evaluated in both aerodynamic and dynamic aspects. The results demonstrate that the circling condition makes a great difference to the performance of rotor vortex, as well as the unsteady aerodynamic loads. With the decrease of the circling radius or the increment of the circling velocity, the thrust of the main rotor increases apparently to balance the inertial force. Meanwhile, the harmonics of aerodynamic loads in rotor disc change severely and an evident aerodynamic load shock appears at high-order components, which further causes a shift-of-peak-phase bending moment in the flap dimension. Moreover, the advancing side of blade experiences second blade/vortex interaction, whose intensity has a distinct enhancement as the circling radius decreases with the motion of vortexes.

It is difficult to simulate the strong interference and serious flow separation of Fenestron by the CFD method based on the widely used RANS equation, and the detailed experimental data, which could be used to validate the aerodynamic and noise numerical methods, is unavailable. The experimental investigation on the aerodynamic and noise characteristics of Fenestron is carried out. In view of the complex internal flow field in duct and the relative motion between the stationary duct and the rotating rotor, a comprehensive aerodynamics and pressure measurement scheme is designed based on the bottom support rig. In this measurement scheme, the thrust generated by the rotating rotor can be measured by the rotating shaft balance, the thrusts from Fenestron are measured by the external balance and the pressures on duct inner wall are monitored by a pressure measuring system. To fully capture the noise directionality of Fenestron, a series of noise observers located at an arc array are arranged. In terms of the Fenestron test models, the baseline model, the performance improvement model based on the high-performance tail rotor and the noise reduction model based on the non-uniform blade distribution are designed respectively. By the designed measurement scheme, aerodynamic forces and pressure distributions and noise were measured for the three different Fenestron models. The results show that the aerodynamic thrusts of the tail rotor and duct increase greatly and the noise increases slightly for the performance improvement model because of the larger aerodynamics. The rotor aerodynamic performance of the noise reduction model is reduced, but the modulation effect of the tail rotor improves the forces of the duct. The noise radiated by the noise reduction model is reduced and a good noise reduction result is obtained in frequency domain.

Numerical simulation methods for unsteady vortex field of helicopter rotor with high resolution and low dissipation TENO8-AA primitive variables reconstruction schemes are established based on moving-embedded grid and Navier-Stokes equations. Firstly, the Targeted Essentially Non-Oscillatory (TENO) scheme are developed by employing ENO-like candidate stencil selection strategy, and the candidate stencil is adopted with optimal weight in smooth region while it is discarded completely in discontinuous region, which reduces the dissipation and dispersion errors and approaches better spectral properties. Then, the aerodynamic characteristics of Helishape-7A model rotor in Blade Vortex Interaction (BVI) state and the flowfield of Lynx rotor in hover are simulated, which validates that the blade tip vortex trajectory with larger wake age and more details of vortex can be captured by TENO8-AA scheme with only a quarter of grid points and half time comparing to WENO-JS scheme. Moreover, the simulation accuracy of thrust coefficient is improved by up to 36%. Finally, the analyses for BVI and aeroacoustic characteristics of Operational Loads Survey (OLS) rotor are conducted, and the different forms of interaction mechanism are explored, such as oblique and parallel interactions. The results indicate that TENO scheme not only ensures the resolution of simulation in discontinuous region, but also minimizes the numerical dissipation in smooth region dominated by blade tip vortex. Therefore, the acoustic pressure peak prediction error of rotor in BVI state is significantly reduced to 5.6% and 0.8% at two microphone locations, respectively.

To study the influence of rotation speed and rotor diameter on the control characteristics and flight performance of a tilt-rotor aircraft, a flight dynamics model and a required power calculation model of the tilt-rotor were established. Trim analysis was carried out under three typical flight modes, and the change of the required power was further analyzed. The results showed that as the rotor speed/diameter increased, the collective pitch and cyclic pitch of the rotor and the pitching angle of the fuselage decreased, and the deflection angle of the elevator increased. As the rotor speed increased, the required power of the tilt-rotor aircraft increased. In helicopter mode, the required power decreased with increasing rotor diameter, while in fixed-wing mode, the required power increased with increasing rotor diameter.

Tilt-rotor UAVs are exceptionally suitable for near-ground unmanned transportation and other fields because of their helicopter mode, transition mode and fixed-wing mode. However, the tilt-rotor transition on trajectory planning presents a challenge, as performance constraints of different modes must be considered and a trajectory planning algorithm is required to optimize a collision-free and short trajectory with low fuel consumption and small cost. To address these challenges, this paper adds performance constraints of the tilt-rotor UAV to the hybrid algorithm by combining the A* algorithm and the artificial potential field method; meanwhile, the rationality of trajectory planning is improved by setting the tilt-rotor transition region separately in the global. Simulation results show that a reasonable trajectory can be planned by setting the tilt-transition region by itself and using a hybrid algorithm.

Electrohydrodynamic (EHD) force produced by corona discharge is considered as a new thrust for solar-powered aircraft and stratosphere balloons in near space. However, its performance at low air pressures remains to be clarified. An experiment of measuring the EHD force at 0.02 atm–1.0 atm (1 atm = 1.01325×10 5 Pa) is carried out with the wire-to-cylinder geometric structure. The ion distribution is analyzed by using the drift-diffusion model with two-dimensional numerical simulation. The experimental result shows that the EHD force is not linearly related to the corona discharge current at low air pressures. Numerical simulation finds that the proportion of ions in the counter-direction electric field increases from approximately 0.36% to 30% when the pressure drops from 1.0 atm to 0.2 atm. As a result, the EHD force with a constant power supply drops faster than the previous theoretical prediction in the ground experiment, suggesting that the consideration of counter-direction EHD force is necessary to improve the ionic wind propulsion efficiency in near-space applications.