To investigate the unsteady aerodynamic characteristics of rotor with Variable Trailing-Edge Camber (VTEC), an unsteady rotor flowfield simulation method is established based on the URANS equation by introducing a deformable moving-embedded grid method. The influence mechanisms of variable-camber amplitude A_m, frequency k, and phase φ₀ on the unsteady aerodynamic characteristics of rotor are analyzed thoroughly, and whereby a VTEC optimization method is proposed through cross-iteration of variable-camber parameters to achieve the dual objectives of hub load suppression and trim maintenance. The numerical experiments indicate that the k = 3 harmonic plays a dominant role in controlling the fluctuation of vertical hub load, which corresponds to the three-bladed rotor. The feasibility of dual load suppression/trim maintenance of VTEC is demonstrated, in which the fluctuation amplitudes of hub loads increments exhibit quasi-linear relationships with A_m for k = 0 - 4 and the fluctuation phase of vertical hub load variation shifts synchronously with the adjustment of φ₀ for k = 2 - 4. The results demonstrate that the proposed load suppression method can effectively decrease the fundamental 3/rev vertical hub load and simultaneously maintain the rotor trim state.

To investigate the aerodynamic interference characteristics between rotor and wing of the tiltrotor aircraft in ground effect, a high-resolution flow field simulation framework was established based on the overset grid technique, third-order MUSCL/CD scheme and RANS/LES (Reynolds-Averaged Navier-Stokes/Large Eddy Simulation) hybrid method. This study focuses on the variations in aerodynamic characteristics of the rotor-wing with different heights above the ground. Results demonstrate that the presence of the ground and wing inhibits the development of rotor downwash, significantly altering the wake evolution process. The radial position of tip vortices exhibits an initial contraction followed by expansion, while the axial migration rate decreases. Furthermore, the helical motion of the blade tip vortices is characterized by phenomena including breakup, pairing and merging. Rotor thrust increases and wing downwash load decreases as ground height decreases. In comparison with the condition without ground effect, rotor thrust and figure of merit increase by up to 6.20% and 8.40% respectively, while wing downwash load drops by 59.05%. A computational method for Lagrangian Coherent Structures (LCS) was developed based on the fourthorder Runge-Kutta time integration method and the fourth-order central difference scheme to investigate the evolution of vortical structures in near-ground flow fields. By computing finite-time Lyapunov exponent fields at different instants, the spatiotemporal evolution of the vortex structures in the rotor-wing flow field under the ground effect was clarified. This revealed the physical mechanism by which recirculating structures maintain stability through continuous transport of fluid parcels, as well as the flow mechanism leading to reduced aerodynamic forces in the blade tip region due to downwash effects.

In order to investigate the dynamic stall characteristics of elliptic airfoils and capture the main unsteady structures of the flow field, a numerical method was established based on unsteady Reynolds-averaged Navier-Stokes equations with SST k-ω and γ-Re_θt turbulence transition models. The method was applied to simulate deep dynamic stall of an elliptic airfoil with a relative thickness of 16%. The unsteady characteristics of the flow field and the variations of aerodynamic coefficient were thoroughly discussed in the time domain. Dynamic mode decomposition (DMD) was employed to extract the characteristics of velocity and pressure fields, followed by flow field reconstruction and error evaluation using a conjugate-mode truncation method based on the theory of modal energy proportion. The results showed that the separation bubbles generated by the elliptic airfoil at the leading edge during the upstroke process served as a precursor to dynamic stall vortex formation. The modes of each order effectively characterized the dynamics of dynamic stall, consistent with the main flow features in the time domain. While conjugate modes contributed to the unsteady components of the flow field, the first-order DMD mode reflected the uniform flow field with characteristics of deep stall of the airfoil. Excluding the first-order mode, the reconstructed flow field retaining 75%—95% energy of the conjugate modes captured variations of the aerodynamic coefficients in the time domain to a certain extent, but failed to describe fine unsteady details accurately. The reconstructed flow field using 99% energy of the conjugate modes demonstrated a decent accuracy in aerodynamic coefficient prediction and effectiveness in dimensionality reduction despite certain deciations during severe flow separation and limited accuracy in capturing fine flow details of near-wall separation bubbles, showing promising potential for engineering applications.

To balance the computational accuracy and efficiency of rotor aeroelastic coupling analysis, the Viscous Vortex Particle Method (VVPM) and the Geometrically Exact Beam Theory (GEBT) tight-coupling method are developed. In aerodynamic aspect, the VVPM method and the Leishman-Beddoes (L-B) dynamic stall model which consider the influence of the elastic deformation are used to simulate the wake viscous flow of the rotor and the unsteady aerodynamic characteristics of blades. In structural aspect, the nonlinear dynamic equations which consider the collective pitch ramp increase have been developed based on the GEBT method. The high efficiency and accuracy characteristics of the coupling method are verified by comparative analyses of the steady/unsteady aerodynamic characteristics in the hovering state. Taking the BO105 rotor as a reference, the elastic responses of the blade and its influence on the aerodynamic characteristics are analyzed during collective pitch ramp increase in the hovering state. The results show that the dynamic stall phenomenon tends to occur when the initial collective pitch is large, leading to a rapid increase of the negative moment, a decrease of the section airfoil normal force coefficient, and a weakening of the thrust overshoot. At the end of the collective pitch ramp increase, the elastic blade has a larger range of thrust overshoot area compared to the rigid blade, resulting in a more serious thrust overshoot phenomenon. The delayed shedding of the rotor wake is captured through rotor vortex field simulation, revealing that the induced velocity of rotor disc is responsible for the thrust overshoot phenomenon of the rotor.

The development of smart materials has made it possible to achieve continuous variable-camber function of airfoil/rotor. To study the influence mechanism of continuous variable-camber airfoil/rotor on the strong unsteady aerodynamic characteristics during dynamic stall, a rotor flowfield simulation method is established based on the moving-embedded mesh method and URANS equation, and the body-fitted mesh of variable-camber airfoil/rotor is reconstructed by the RBF mesh deformation method. The effectiveness of the mesh method and the numerical simulation method is verified by comparing the unsteady aerodynamic characteristics of the NACA0012 airfoil, the SA349/2 rotor and the SMART rotor with trailing edge winglet. A comparison with the traditional rigid trailing edge winglet reveal the principle and advantage of continuous variable-camber technology for improving the average lift-to-drag ratio of airfoil. The complex effect of variable-camber airfoil/rotor on dynamic stall characteristics is captured, and the influences of motion frequency, amplitude and initial phase of the variable-camber on the unsteady aerodynamic characteristics of the airfoil, as well as the unsteady aerodynamic load variation characteristics of the three-dimensional rotating blade with different azimuth angles and blade sections, are fully considered. The results demonstrate that the continuous variable-camber technology can effectively control the aerodynamic performance of the rotor through collaborative allocation of variable-camber frequencies. An approximate linear relationship between the variable-camber amplitude and the unsteady aerodynamic increment of the airfoil/rotor is observed even under the condition of dynamic stall. The initial phase of variable-camber should be consistent with the cyclic pitch control of rotor to suppress the divergence of sectional aerodynamic coefficient. The above rules indicate that the continuous variable-camber technology has great potential in rotor trim or vibration reduction.

To analyze the aeroelasticity of Trailing Edge Flap (TEF) rotors in forward flight, a high-precision analysis method based on Computational Fluid Dynamics/Computational Structural Dynamics (CFD/CSD) coupling is established. A set of moving-embedded grid method for rotors with TEF is developed by the parametric method. In terms of aerodynamic analysis, the high-precision CFD method, based on the traditional aerodynamic model, is introduced to accurately simulate the unsteady flow field and aerodynamic characteristics of the rotor. In terms of structure, the dynamic model of rigid TEFs is established, the influence of the rigid TEF motion on the rotor system is considered in the form of additional mass, damping, stiffness and other matrices, and the nonlinear dynamic equation of the rotor system is derived based on Hamilton's principle and Timoshenko beam model. An elastic grid deformation method based on algebraic transformation is adopted, and the unsteady fluid-solid coupling strategy is developed. Results comparison of the aeroelastic loads of the model rotor and the SA349/2 rotor verifies the effectiveness of the CFD/CSD coupling method. On this basis, taking the SMART rotor with TEF as an example, the calculation results are compared with the equivalent normal force coefficient of the blade sections. Emphasis is placed on the analysis of the Fulton model rotor, studying the variation of blade root torque in hover and flap moment amplitude in forward flight, and comparing them with literature results. Results show that the proposed CFD/CSD coupling method can improve the analysis accuracy of the unsteady aeroelastic load of the rotor, and accurately capture the Blade-Vortex Interaction (BVI) phenomenon of the low-speed forward flight rotor, with the average error of flap bending moment controlled at 11.68%. Meanwhile, the natural frequency error of the blade with TEF is smaller than 4.0%, and the average error of flap bending moment amplitude of the blade root is 15.15%, meaning that the aeroelastic characteristics of the TEF can be effectively simulated. Under the single TEF control law, the hub vertical force F_z of the Fulton rotor can be effectively controlled, and the load amplitude is reduced by more than 80%.

To investigate the influences of interference between the fuselage/propeller on the aerodynamic and aeroacoustic source characteristics in high-speed, firstly, a highly-robust embedded grid method was established in which hybrid grids of fuselage non-structure/propeller structure were adopted. An outer boundary identification method was proposed by adapting the outer boundary to the solid object so as to circumvent the space overlap between the grid domain and the other solid object. Then the Computational Fluid Dynamics (CFD) method based on the Unsteady Reynolds Averaged Navier-Stockes (URANS) equations was established for the simulation of helicopter fuselage/propeller flowfield. Furthermore, the methods based on the Ffowcs Williams and Hawkings (FW-H) equations were applied for fuselage and propeller noise prediction. The effectiveness of the developed method was verified by comparing the simulation results with the results of the NASA ROBIN/rotor experiment and the AH-1G noise experiment. Finally, the flowfield of the fuselage/propeller was simulated for the analysis on the interference mechanisms and aeroacoustics characteristics, and some meaningful conclusions were obtained. Under the interference of the fuselage, a fluctuation of about 5.34% occurred on the propeller thrust while the fluctuation of the single blade thrust was more significant. The Sound Pressure Level (SPL) of the propeller under interferences increased significantly (compared to the isolated propeller). The increased values were positively correlated with the advance ratio. In addition, the frequency domain characteristics of the total noise were altered due to the loading noise induced by the fluctuation of fuselage surface pressure.

The coaxial rigid rotor high-speed helicopter based on the forward blade concept has a high maximum flight speed, but the cooperation of the coaxial rigid rotor and the propellor brings the problem of handling redundancy. To improve the flight performance of the high-speed helicopter, this paper carries out the analysis of the influence of the rotor/thrust blade control strategy. Firstly, a high-precision Computational Fluid Dynamics (CFD) method for predicting the aerodynamic forces of coaxial rigid rotor, propellor and fuselage is established. Secondly, a robust trim method for redundant control of coaxial rigid rotor/propellor is proposed based on CFD-Surrogate Model (SM) -Genetic Algorithm (GA). On this basis, the influence of control strategies such as rotor/propellor forward thrust distribution, rotor lift offset and fuselage attitude on the flight performance of high-speed helicopter is analyzed, and the control strategies that can effectively improve the maximum forward flight speed and practical lift limit of high-speed helicopter are obtained. The results show that with fixed fuselage pitch angle, the maximum horizontal flight speed of the high-speed helicopter can be increased by 10.78% when a small part of the forward thrust is distributed by the rotor. When flying at sea level, the helicopter has the maximum forward speed when the lift offset of the rotor is 0.3, and has the maximum service ceiling when the lift offset is 0.2. When the fuselage pitch angle is -1°, the high-speed helicopter has the maximum forward flight speed and service ceiling. If the fuselage pitch angle is too large or too small, the forward flight performance will be reduced.

To grasp rotor aeroacoustic radiation characteristics during helicopter flight close to the ground, firstly, based on the compressible Reynolds average Navier-Stokes equations and the Ffowcs Williams-Hawkings equations, a set of aerodynamic and aeroacoustic analysis method suitable for rotor/fuselage interaction in ground effect state was developed. The reliability of the established method was verified by comparing the ground effect test of Lynx tail rotor and the NASA rotor/fuselage interaction test with calculation results. Then, the influence of fuselage and hub on the rotor flowfield was investigated. It was found that the fuselage could change the rotor aerodynamic load distribution and increase the aeroacoustic radiation level of the intermediate order (8th–12th order). The aeroacoustic contribution of the blade middle section (r=0.4R–0.7R, R is the radius of rotor) could be weaken by the hub. Finally, the aerodynamic and aeroacoustic characteristics of the rotor at different heights (h) above the ground were investigated and a “critical ground height” with Blade/Vortex Interaction (BVI) noise was found. The results show that the rotor thrust and fuselage lift decrease with the increase of ground height. When h>1.8R, the influence of ground effect can be ignored. When h=0.6R, the inflow of the rotor will be changed by the blocking effect of the ground and the fuselage. Especially, load fluctuations are induced severely in the blade tip area and obvious BVI noise is received at the characteristic observation points. According to the time/frequency characteristics, this paper suggests a “critical ground height” of approximately 0.7R, which provides a reference for the noise suppression when the helicopter flies close to the ground.

By combining the characteristics of high hovering performance and control efficiency of coaxial-rotor helicopters and the advantage of high-speed forward flight performance of double-thrust-propeller compound configuration helicopters, the coaxial rotor was applied to the double-thrust-propeller compound configuration helicopter. The single rotor was changed to a coaxial rotor based on the configuration of the X3 compound helicopter. To study the influence of coaxial rotors on the aerodynamic characteristics of compound configuration helicopters, a fast trim method was established. On this basis, the aerodynamic characteristics of different configuration helicopters were analyzed. The results indicate that the aerodynamic characteristics of the coaxial-rotor helicopter have good symmetry compared with the single-rotor configuration. In the case of maintaining good high-speed forward flight performance, the coaxial-rotor configuration can significantly improve the hovering and low-speed flight performance. When the hovering efficiency is increased by 6.8%, and the forward flight speed is 100 km/h, the total required power is reduced by 23.1%. The aerodynamic interference of the coaxial rotor to the propeller at low speeds is greatly lower than that of the single rotor.