Rotor In Forward Flight Research Articles

Wind Tunnel Investigation of the Icing of a Drone Rotor in Forward Flight

The Bell Textron APT70 is a UAV concept developed for last mile delivery and other usual applications. It performs vertical takeoff and transition into aircraft mode for forward flight. It includes four rotor each with four rotating blades. A test campaign has been performed to study the effects of ice accretion on rotor performance through a parametric study of different parameters, namely MVD, LWC, rotor speed, and pitch angle. This paper presents the last experimentations of this campaign for the drone rotor operating in forward flight under simulated icing conditions in a refrigerated, closed-loop wind tunnel. Results demonstrated that the different parameters studied greatly impacted the collection efficiency of the blades and thus, the resulting ice accretion. Smaller droplets were more easily influenced by the streamlines around the rotating blades, resulting in less droplets impacting the surface and thus slower ice accumulations. Higher rotation speeds and pitch angles generated more energetic streamlines, which again transported more droplets around the airfoils instead of them impacting on the surface, which also led to slower accumulation. Slower ice accumulation resulted in slower thrust losses, since the loss in performances can be directly linked to the amount of ice accreted. This research has not only allowed the obtainment of very insightful results on the effect of each test parameter on the ice accumulation, but it has also conducted the development of a unique test bench for UAV propellers. The new circular test sections along with the new instrumentation installed in and around the tunnel will allow the laboratory to be able to generate icing on various type of UAV in forward flight under representative atmospheric conditions.

Analytic Prediction of Rotor Broadband Noise with Serrated Trailing Edges with Applications to Urban Air Mobility Aircraft

Theoretical predictions for broadband noise generated by rotors with serrated trailing edges are presented. This study extends Lyu and Ayton’s model for serrated trailing-edge noise to account for factors such as finite span length, the accurate acoustic attenuation rate of trailing-edge noise, modified spanwise wavenumber, and the proper scattering factor. This new model is implemented in the rotorcraft broadband noise tool, UCD‐QuietFly. The validation results demonstrate a satisfactory level of agreement with the experimental data for serrated trailing‐edge noise of a wing section, a hovering rotor, and a forward‐flight rotor. Next, the effects of serration parameters, such as the number of serrations, serrations starting radial location, serration geometry, and serration height, on rotor broadband noise are investigated. For example, it is observed that sine-wave, chopped peak, or sawtooth serration shapes provide the most significant noise reduction among the various shapes considered. Finally, broadband noise reduction through serrations is applied to two urban air mobility aircraft scenarios: a six‐passenger quadrotor and a quiet helicopter. For the quadrotor, serrated trailing edges show a noise reduction potential of over 9 dB, with higher reductions observed at higher tip speeds and fewer blades. A quiet helicopter equipped with serrated blades in forward flight achieves a maximum noise reduction of over 10 dB in the forward direction.

Numerical investigation on the use of various blade tips for the helicopter's main rotor blade in forward flight regarding aerodynamic performance and HSI noise considerations

The goal of this research is to investigate aerodynamic performance and high-speed impulsive (HSI) noise of various blade tips for the helicopter's main rotor in forward flight. For this purpose, three-dimensional compressible flow field around the rotor blades is numerically simulated by the solution of the unsteady Reynolds averaged Navier-Stokes (URANS) equations with the "SST k-ω" turbulence model. Solution is quantified by the calculation of torque, drag, and thrust coefficients of CQ, CD, CT, and L/D ratio. HSI noise on the rotor plane is assessed by the calculation of sound pressure level at certain points of the flow field using the Ffowcs Williams-Hawking (FW-H) equation. Blade tips are eleven and include anhedral, dihedral, Eagle and combinations of them, and also Blue Edge and rectangular. The combinations are inspired by the tip configurations of the fixed wings. In this regard effects of the tip geometry parameters including curvature and height of its parts have been considered as well. All of the tips are examined for two flight conditions of Caradonna and Tung two-blade rotor with fixed pitch angle and time-varying pitch angles. Based on the present results, it is found that blade tips of dihedral family do not improve the aerodynamic performance of the helicopter blade. Also, addition of anhedral part to the blade tip improves the L/D ratio. According to the present acoustic analysis, it was seen that after the Blue Edge, the Eagle tip stands in order of priority. For the condition of Caradonna and Tung two-blade rotor with tip Mach number of 0.8 and advance ratio of 0.2, HSI noise of the Eagle-tip blade decreases by 2.39 % compared to that of the two-blade rotor with rectangular tip. However, the aerodynamics performance of the Eagle tip is much better than that of the Blue Edge.

Analysis of the Aeroacoustic Characteristics of a Rigid Coaxial Rotor in Forward Flight Based on the CFD/VVPM Hybrid Method

This study develops a hybrid solver with reversed overset assembly technology (ROAT), a viscous vortex particle method (VVPM), and a CFD program based on the URNS method, in order to study the aerodynamic and acoustic characteristics of coaxial rigid rotors. The aerodynamic load of the “AH-1G” helicopter rotor is first calculated based on the hybrid method and compared with available experimental data. The prediction of the linear noise of the OLS rotor is then performed and the obtained results are compared with available experimental data. These results allow the evaluation of the accuracy of the hybrid method for emulating rotor aerodynamics and acoustics. Afterwards, the hybrid and CFD methods are applied to obtain the aerodynamic and acoustic characteristics of the given coaxial rigid rotor model, while taking into account the trim of the collective pitch. The obtained results demonstrate that the hybrid method has high proficiency in capturing blade–vortex-interaction impulsive loads and high computational efficiency in predicting associated loading noise characteristics. Furthermore, the effect of the hybrid method on the noise characteristics of coaxial rigid rotors under a different advance ratio, blade tip speed, shaft angle, and other conditions, as well as the impact of the upper and lower rotors on the noise contribution of the coaxial rotor are analyzed. Finally, the impacts of the initial phase and the vertical spacing on the sound pressure level are studied.

Geometrically exact aeroelastic stability analysis of helicopter composite rotor blades in forward flight

In this paper, a sophisticated method for analyzing the geometrically exact aeroelastic stability of composite rotor blades in forward flight has been proposed. The unique advantage of the proposed method is to use the geometrically exact beam model with the updated Variational Asymptotic Beam Sectional analysis (VABS) for structural modeling to not only calculate the blade cross-sectional properties and blade spanwise deformations more accurately, but also take into account the effect of transverse shear deformation and blade initial twist and curvatures sophisticatedly. The Peters finite state airloads theory and the Peters-He finite state dynamic inflow model are coupled to calculate the three-dimensional unsteady airloads applied on the blade. The auto-pilot trim scheme is adopted to determine the blade pitch controls. The finite element spatial discretization method, the Newmark numerical integration method, the Newton-Raphson method and the moving-block analysis method are combined to solve the established geometrically exact aeroelastic equations and obtain the geometrically exact aeroelastic stability of composite rotor blades. The proposed method is validated by correlation with existing analytical results. The investigations show that the effect of transverse shear deformation on the aeroelastic stability of composite hingeless rotors in forward flight is non-negligible. Blade initial curvatures have significant effect on the aeroelastic stability of composite hingeless rotors in forward flight, which is not only related to the blade elastic couplings, but also to the advance ratio.

Implementation and Linearization of State-Space Free-Vortex Wake Models for Rotary- and Flapping-Wing Vehicles

This article describes the implementation and linearization of free-vortex wake models in state-variable form as applied to rotary- and flapping-wing vehicles. More specifically, the wake models are implemented and tested for a UH-60 rotor in forward flight and for a hovering insect representative of a hawk moth. A periodic solution to each wake model is found by time marching the coupled rotor/wing and vortex wake dynamics. Next, linearized harmonic decomposition models are obtained and validated against nonlinear simulations. Order reduction methods are explored to guide the development of linearized wake models that provide increased runtime performance compared to the nonlinear and linearized harmonic decomposition wake models while guaranteeing satisfactory prediction of the periodic response of the wake. This constitutes a first attempt to extend free-vortex wake methods in state-variable form, originally developed for rotary-wing applications, to flapping-wing flight.

A data-driven reduced-order model for rotor optimization

Abstract. For rotor design applications, such as wind turbine rotors or urban air mobility (UAM) rotorcraft and flying-car design, there is a significant challenge in quickly and accurately modeling rotors operating in complex, turbulent flow fields. One potential path for deriving reasonably accurate but low-cost rotor performance predictions is available through the application of data-driven surrogate modeling. In this study, an initial investigation is undertaken to apply a proper orthogonal decomposition (POD)-based reduced-order model (ROM) for predicting rotor distributed loads. The POD ROM was derived based on computational fluid dynamics (CFD) results and utilized to produce distributed-pressure predictions on rotor blades subjected to topology change due to variations in the twist and taper ratio. Rotor twist, θ, was varied between 0, 10, 20, and 30∘, while the taper ratio, λ, was varied as 1.0, 0.9, 0.8, and 0.7. For a demonstration of the approach, all rotors consisted of a single blade. The POD ROM was validated for three operation cases: a high-pitch or a high-thrust rotor in hover, a low-pitch or a low-thrust rotor in hover, and a rotor in forward flight at a low speed resembling wind turbine operation with wind shear. Results showed that reasonably accurate distributed-load predictions could be achieved and the resulting surrogate model can predict loads at a minimal computational cost. The computational cost for the hovering blade surface pressure prediction was reduced from 12 h on 440 cores required for CFD to a fraction of a second on a single core required for POD. For rotors in forward flight, cost was reduced from 20 h on 440 cores to less than a second on a single core. The POD ROM was used to carry out a design optimization of the rotor such that the figure of merit was maximized for hovering-rotor cases and the lift-to-drag effective ratio was maximized in forward flight.

Experimental Investigations on Flow Control of the Rotor via the Synthetic Jets in Forward Flight

To study the effects of synthetic jet control on the aerodynamic performance of a rotor in forward flight, we conducted a series of experiments with varying rotor rotation speeds and free stream velocities. In the test, we used a six-component balance and a PIV system and designed a blade with a particular structure that covered the frame. The experimental results revealed that the synthetic jet could effectively delay flow separation over the blade and enhance the aerodynamic efficiency of the rotor. Moreover, we investigated how different jet parameters influenced the flow control effects of synthetic jets on the rotor’s aerodynamic characteristics. We drew some valuable conclusions from our analysis. In forward flight, the jet located closer to the leading edge of the blade had a stronger impact on improving the aerodynamic performance of the rotor. The jet with a 90° jet angle increased the rotor normal force by 225%, which was the maximum possible increase, while the jet with a 30° inclined angle had the best control effects on preventing flow separation in the retreating blade. Our study provides valuable insights into the use of synthetic jets for rotor flow control and suggests possible applications for improving rotorcraft performance and stability.

Pseudospectral-Method-Based Framework for Rotorcraft Stability and Trim Analysis

The paper demonstrates applications of pseudospectral-method-based framework to rotorcraft dynamics. The framework is developed to incorporate blade-to-blade symmetry (fast-Floquet) to reduce computational time. The methodology is applied to compute Floquet exponents of an isolated rotor in forward flight using the periodic eigenvector equations, and computational performance is assessed. The framework is also applied to a trim problem that determines time-varying control input for equilibrium, as well as meets stability criteria in the form of minimum damping of the system at a prescribed value. Through the applications considered, the paper brings out the advantages of the use of the pseudospectral framework, particularly its amenability to problems that involve optimization.

Actuator Disk Model with Improved Tip Loss Correction for Hover and Forward Flight Rotor Analysis

A novel actuator disk model (ADM) coupled with lifting-line theory is proposed in this paper. Several virtual planform blades are placed on a disk plane with a constant azimuthal interval, and the lifting-line theory is applied to each blade to predict the effective angle of attack. The proposed model considers the local lift and drag forces acting on disk surface cells by interpolating the predicted effective angle of attack with various azimuth angles to the actuator disk plane; therefore, the proposed model considers individual blade tip vortices without tip loss functions. Experimental data for hover and forward flight rotors are used to validate the proposed model. For hovering flight, sectional thrust based on collective pitch angles predicted by the modified ADM was similar to that obtained in the experiments. For forward flight, the inflow above the rotor estimated by the proposed ADM was similar to that obtained in the experiments and by using other numerical methods. Thus, the developed ADM can be used for rotor performance analysis under the main flight conditions of V/STOL.

CFD analysis on the performance of a coaxial rotor with lift offset at high advance ratios

The aerodynamic performance of an isolated coaxial rotor in forward flight is analyzed by a high-fidelity computational fluid dynamics (CFD) approach. The analysis focuses on the high-speed forward flight with an advance ratio of 0.5 or higher, which is the ratio of the forward speed to the rotor tip speed. The effect of the degree of the rolling moment on the rotor thrust, called lift offset, is studied in detail. The coaxial rotor model is a pair of contrarotating rotors, each rotor consisting of two untwisted blades with a radius of 1.016 m. The pitch angle of the blades is controlled by both collective and cyclic as in a conventional single main-rotor helicopter. CFD analysis is performed using a flow solver based on the compressible Navier-Stokes equations with a Reynolds-averaged turbulence model. Laminar/turbulent transition in the boundary layer is taken into account in the calculation. The rotor trim for target forces and moments is achieved using a gradient-based delta-form blade pitch angle adjusting technique in conjunction with CFD analysis. The reliability of the calculations is confirmed by comparison with published wind tunnel experiments and two comprehensive analyses. Applying the lift offset improves the lift-to-effective drag ratio (lift-drag ratio) and reduces thrust fluctuations. However, in the case where the advance ratio exceeds 0.6, the lift-drag ratio drops significantly even if the lift offset is 0.3. The thrust fluctuation also increases with such a high advance ratio. Detailed analysis reveals that the degradation of aerodynamic performance and vibratory aerodynamic loads is closely related to the pitch angle control to compensate for the reduction in thrust on the retreating side due to the increased reverse flow region. It is effective to reduce the collective and longitudinal cyclic pitch angles for the improvement of the aerodynamic performance of coaxial rotors with an appropriate lift offset.

Explicit Source 기반의 Actuator Model 사용을 위한 OpenFOAM의 중첩 격자 라이브러리 수정

This study purpose is modifying the overset grid library of OpenFOAM to use an explicit source in the overset grid system. To this end, the functions of the OpenFOAM"s overset grid library that classify the computational volume cells was modified. Actuator Disk Model based on an explicit source was employed to confirm the functionality of the modified functions. The performance analysis at hover flight and the inflow analysis at forward flight were performed using the modified overset library. Since the hovering rotor performance results with single and overset meshes were almost similar, it was confirmed that the modified overset library well performed at hover flight. Through the forward flight rotor analysis, it was confirmed that the rotor inflow analysis results were similarly predicted in the single and overset meshes. Through previous analysis results, it was confirmed that the modified overset library worked properly in the problem of using explicit source.

Numerical Investigation on Aerodynamic Performance and Interaction of a Lift-Offset Coaxial Rotor in Forward Flight

Numerical Investigation on Aerodynamic Performance and Interaction of a Lift-Offset Coaxial Rotor in Forward Flight

Numerical analysis of rotor aeroacoustic characteristics during collective pitch aperiodic variation in hover

Based on the FW-H equations and the CFD method, the rotor aeroacoustic characteristics during collective pitch aperiodic variation in hover are calculated and analyzed. First, a set of analysis method for aperiodic rotor aeroacoustic characteristics is developed. The aerodynamic and aeroacoustic characteristics of the BO-105 rotor in hover, the AH-1G rotor in forward flight and the NACA rotor in a ramp increase of collective pitch are calculated, and the employed numerical analysis method is validated through comparisons with experimental data. Then, the aeroacoustic characteristics of the NACA rotor during an oscillating collective pitch are analyzed, and the sound pressure peak at different collective pitch is discussed in detail. Finally, the BO-105 rotor in hover during collective pitch aperiodic variation is calculated. In addition, parameters, such as the collective pitch change rate, are quantified, and some conclusions are obtained. The configuration of collective pitch change rate will influence the directivity of rotor noise in the collective pitch increasing process, and the loading noise will be smaller if the change rate is decreasing.

Transient Thrust Analysis of Rigid Rotors in Forward Flight

The purpose of this study was to investigate and quantify the transient thrust response of two small rigid rotors in forward flight. This was accomplished using a distributed doublet-based potential flow method, which was validated against wind-tunnel experimentation and a transient CFD analysis. The investigation showed that for both rotors, advancing and retreating blade effects were predicted to contribute to transient thrust amplitudes of 5–30% of the mean rotor thrust. The thrust output amplitudes of individual rotors blades were observed to be 15–45% of the mean rotor thrust, indicating that it is not uncommon for the thrust output variation of an individual rotor blade to approach the same value as the mean thrust output of the rotor itself. In addition to this, the theoretical analysis also illustrated that higher blade thrust oscillations resulted in pronounced asymmetric rotor wake structures.

Validation and insight of a full-scale S-76 helicopter rotor using the Lattice-Boltzmann Method

This paper describes scale-resolving unsteady flow simulations performed using a Lattice-Boltzmann based solver, applied, for the first time, to a full-scale helicopter main rotor in forward flight. The simulation setup is obtained through a fully automated in-house developed workflow. Validation is performed by comparing thrust and power coefficients with reference wind tunnel data; with and without trimming based on an in-house developed Newton-Raphson trim algorithm. Results for different rotor advance ratios yield insight in both aerodynamic and acoustic quantities for the rotor, highlighting both mean as well as unsteady behavior. Finally, digital helicopter noise certification is performed by calculating the ground noise map for a helicopter in landing condition.

Research on the forward flight performance of rotor based on variable-droop leading edge

Aiming at the dynamic stall phenomenon of the retreating side of the rotor in forward flight, the existing flow control method of dynamic leading edge droop was applied to the flow control of forward-flying rotor at three-dimensional scale. A numerical simulation method based on variable droop leading edge is established in this paper. The seesaw rotor is taken as the research object, the moving overset mesh method and RBF grid deformation technology are used, the integral form of Reynolds average N-S equation is the main control equation. The influence of the dynamic leading edge at r/R=0.75~1 on the aerodynamic characteristics of the rotor when the forward ratio is 0.3 is investigated. It is found that variable droop leading edge on the retreating side can effectively inhibit the generation and development of separation vortices near the trailing edge, and has a significant effect on lifting lift coefficient and section normal force coefficient, reducing torque coefficient, and thus improving the equivalent lift-drag ratio of the rotor. In a certain range, the control effect is better with the increase of the droop amplitude under the leading edge.

Experimental investigation on the performance degradation of helicopter rotor due to ice accretion

Helicopter rotor performance can be significantly altered by ice accretion which directly degrades the handling qualities and safety during flight. To obtain understanding into the effect of ice accretion on a helicopter rotor in forward flight under different icing environments, an experimental campaign was conducted in an icing wind tunnel using a rotor model with diameter at 2 m. Research was emphasized on revealing effect of liquid water content and static temperature on the rotor performance degradation. The dynamic thrust and torque of the rotor were measured by a six-component balance. Ice shape of the airfoils at specific rotor blade spanwise locations was obtained through the ice cutting process; meanwhile, the entire ice topology on the blade was scanned by means of the three-dimensional scanning technique. Results showed that ice was mainly accreted on the leading edge and lower surface of the blade. The static temperature of a icing cloud has a noticeable effect on the regime of accreted ice. The ice formation was transparent glaze at −8°C and transformed to milky rime when changed the static temperature to −20°C. Comparing to rime ice, the glaze ice on the blade is more influential to the rotor performance degradation. Particular attention is given to the fact that under extreme icing conditions, over 30% degradation of transient rotor thrust was observed.

Some improvements of hybrid trim method for a helicopter rotor in forward flight

The main difficulty of the trimming for helicopter rotor CFD (Computational Fluid Dynamics) lies in the high computational cost and low computational efficiency. The BET/CFD hybrid trim method, which uses the low-fidelity BET (Blade Element Theory) to calculate the Jacobian matrix and high-fidelity CFD method to get the aerodynamic force and moment for comparison with the target values, has been widely used in industry due to its high efficiency and low computational cost compared with the pure CFD method. However, this method also has two disadvantages: a complex C81 table needs to be established before the BET works; because of the low accuracy in computing the Jacobian matrix, the number of trim iterations is too large and the trimming may not even converge. In this paper, two improvements of the hybrid trim method have been proposed. Firstly, an Artificial Neural Networks (ANN) model has been developed to predict the aerodynamic coefficients in the BET. The ANN model greatly improves the efficiency and works for all kinds of airfoils without any extra operation. The second improvement is to divide the trimming process into two stages: the preliminary stage in which the Jacobian matrix is calculated using the BET with an ANN model, and the accurate stage in which the Jacobian matrix is obtained by the CFD method. The two-stage method avoids the second shortcoming mentioned above. The AH-1G rotor and ONERA 7A rotor in forward flight have been investigated to validate the present method. The trim results agree well with the flight test results and other computational results.

Actuator Surface Model with Computational-Fluid-Dynamics-Convected Wake Model for Rotorcraft Applications

An actuator surface model has been developed that couples an aerodynamics model to a computational fluid dynamics solver. The model is designed for simulation of rotors in complex operating environments, and it advances a recently developed actuator surface model through the addition of two new elements: a computational fluid dynamics (CFD)-convected wake model, and an improved coupling algorithm. The wake model uses Lagrangian particle tracking to determine the geometry of the rotor wake from the CFD solution. Prior study of the influence of the maximum wake age in the model led to the conclusion that only 45 deg of wake is required for accurate resolution of air loads. The S-76 rotor in hover and the UH-60A rotor in forward flight are used for validation. Thrust coefficients, torque coefficients, and figures of merit over a range of collective settings agree well with experimental data for the S-76 rotor. Sectional normal force coefficients in forward flight are compared with experimental data, results from resolved-blade simulations, and other rotor models: again showing good agreement. These results suggest the new actuator surface model can predict time-accurate rotor loads where accurate resolution of the near-blade flowfield and high-frequency loads, such as blade–vortex interactions, are not critical.