Blade-vortex Interaction Research Articles

Computational Investigation on Unsteady Loads of High-Speed Rigid Coaxial Rotor with High-Efficient Trim Model

A computational fluid dynamics method is built to study the unsteady aerodynamic loads of a high-speed rigid coaxial rotor model, taking account of lift offset (LOS). The flowfield is simulated by solving Reynolds Averaged Navier–Stokes equations, and moving overset mesh is adopted to include blade motions. A high-efficient trim model for coaxial rotor is developed, where the “delta method” is implemented. Performance of Harrington rotor-1 is calculated for validation. Forward flight cases in three advance ratios are conducted. Results indicate that the temporal thrusts of coaxial rotor at low advance ratio share some fluctuations similar to hover state. In forward flight, the impulsive thrust fluctuations caused by blade-meeting are obviously exhibited around 270° for upper blades, and the strengths increase with the increase of LOS and advance ratio. At higher advance ratios, the blade thrusts of the upper and lower rotors tend to be the same. At the advance ratio of 0.6, two new kinds of Blade–vortex interaction (BVI) are captured. One is the parallel BVI caused by the root vortex and the other is the complex interaction among the tip vortex, root vortex and the rear blade.

Electrically Controlled Rotor Blade Vortex Interaction Airloads and Noise Analysis Using Viscous Vortex Particle Method

An electrically controlled rotor (ECR), also called a swashplateless rotor, replaces a swashplate with a trailing‐edge flap system to implement primary rotor control. To investigate the aerodynamic characteristics of an ECR in blade‐vortex interaction (BVI) condition, an analysis model based on the viscous vortex particle method, ECR blade pitch equation, and the Weissinger‐L lifting surface model is established. In this model, the ECR wake flow field vorticity is discretized as multiple vortex particles, and the vorticity‐velocity form of the Navier‐Stokes equation is solved to simulate the transport diffusion of the vorticity. The flap motion‐inducing blade‐pitch movement is obtained by solving the ECR blade‐pitch movement equation via the Runge–Kutta fourth‐order method. On the basis, BVI noise radiation of an ECR is evaluated using the Ffowcs Williams and Hawkings (FW‐H) equation. Based on the present prediction model, the aerodynamic and acoustic characteristics of a sample ECR in BVI condition are analyzed. The results show that since the BVI event of the ECR on the advancing side is mainly caused by the interaction between the flap tip vortex and the blade, the blade spanwise range of ECR BVI occurrence on the advancing side is smaller than that of the conventional rotor. In addition, the magnitude of the maximum sound pressure level on the advancing side as well as on the retreating side of the ECR is also different from that of the conventional rotor, which is consistent with the difference in the airloads between the ECR and conventional rotor. Furthermore, a study was performed to examine the effect of the pre‐index angle on the BVI‐induced airloads and noise. The amplitude of the impulsive airloads of the ECR on the advancing side is increased with the increase in pre‐index angle, while the amplitude of the impulsive airloads of the ECR on the retreating side is decreased. Indeed, when the pre‐index angle of the sample ECR is 8 degrees, the retreating‐side noise radiation lobe is almost disappeared. In addition, the different intensity of wake vorticity is the main reason for the differences of the BVI‐induced airloads and noise among the ECR with different pre‐index angles.

A fully-coupled CFD/CSD computational approach for aeroelastic studies of helicopter blade-vortex interaction

Blade-vortex interaction (BVI) is one of the main sources of noise and vibrations in helicopters. The aeroelastic response of the blade, to BVI phenomenon, is one of the main factors affecting the helicopter's aerodynamic performance and stability. In the present research we developed and implemented a novel CFD-based, strong (two-way) time-domain coupling, mathematical model, for the numerical prediction of the aeroelastic response of the helicopter blade to BVI. Another novelty, of the fully coupled CFD/CSD approach, is that the fluid flow is solved using the large-eddy simulation (LES) approach. In the present research, the aeroelastic response of a symmetric NACA0012 airfoil to BVI is computed using a two-degree of freedom (2-DOF) mass/spring model. The numerical studies show that the blade-vortex offset distance, (h), plays a key role in the mechanism of BVI and therefore, it influences the aerodynamic coefficients. The computational studies show that the BVI phenomenon diminishes with the increase of the offset distance (h). Also, it was observed that the elastic nature of the airfoil acts as a damper and it diminishes the effect of BVI on the aerodynamic coefficients. Therefore, for a flexible airfoil, the lift coefficient exhibits a decay, at the instant of blade-vortex interaction, when compared with rigid airfoil.

The effect of aerofoil camber on cycloidal propellers

PurposeThis paper aims to investigate the impact of aerofoil camber on the performance of micro-air-vehicle-scale cycloidal propellers.Design/methodology/approachFirst, experiments were conducted to validate the numerical methodology. After that, three turbulent models were compared to select the most accurate one. Then, 2D numerical simulation was carried out on 11 aerofoils with different cambers, including five cambered aerofoils, one symmetrical aerofoil and five inverse cambered aerofoils. The inverse cambered aerofoils are symmetrical about the chord line to the corresponding cambered ones.FindingsThe cycloidal propeller with large cambered aerofoil gives the lowest hovering efficiency, but with symmetrical aerofoil or small inverse cambered aerofoil shows the highest. Also, blades with large cambered aerofoil display high performance at the upper part of its trajectory, while with symmetrical aerofoil or the inverse cambered aerofoil have their best at the lower part. In addition, intensified downwash can be observed in the rotor cage for all cases. When a blade runs through the top-left part of its circle path, all cases display the feature of deep dynamic stall. When the blade travels through the nadir of its path, the actual angle of attack is close to zero due to the strong downwash. Furthermore, there exits intensified blade-vortex interaction induced by the preceding blade for large cambered aerofoils at the lower-right part of its trajectory.Practical implicationsThis paper develops a new cycloidal propeller which is more efficient than the one already present.Originality/valueThis paper discovers that the aerofoil camber is a vital design parameter in the performance of cycloidal propeller, and the authors expect that the rotor with deformable aerofoil on camber would achieve much higher efficiency.

New Approach for Aerodynamic and Aeroacoustic Analysis of Actively Controlled Flaps Rotor

A new approach for aerodynamic and aeroacoustic analysis of actively controlled flaps (ACF) rotor is established to predict airloads and noise reduction. The approach is divided into three parts. The first is an unsteady aerodynamic model of the ACF described through a panel method, and the unsteady pressure under interaction of rotor tip vortex and ACF trailing-edge vortex, which are depicted through a viscous vortex particle method, are taken into account. The second is the influence of the ACF on trajectory of the wake and ACF trailing-edge vortex, which is carried out by a new imaged vortex model to capture blade-vortex interaction (BVI). The third is an aeroacoustic model coupled into the aerodynamic model by transmitting the unsteady pressure to Ffowcs Williams and Hawkings (FW-H) equation. The present approach is applied to AH-1G rotor, and the results show that the predicted unsteady airloads, high-speed impulsive (HSI), and BVI noise are compared favorably with experiment and Reynolds-averaged Navier–Stokes (RANS)/FW-H results. Finally, the variation of airload and noise reduction of ONERA Active Blade Concept rotor (ABC) is predicted and compared with experiment. It is shown that the present approach predicts the airload and noise reduction of the ACF with different phases and flap locations with good accuracy.

Blade-Vortex Interaction Noise Controller Based on Miniature Trailing Edge Effectors

The present work focuses on the alleviation of Blade Vortex Interaction (BVI) noise annoyance through a control methodology generating high-frequency aerodynamic BVI counter-actions. The low-power requirements make the Micro-Trailing Edge Effectors (MiTEs) particularly suited for this kind of application. The controller layout is set by observing the BVI scenario while the actuation law is efficiently synthesized through a process based on an analytical unsteady sectional aerodynamic formulation. The validation of the proposed control methodology is carried out through numerical investigations of a realistic helicopter main rotor in flight descent, obtained using computational tools for potential-flow aerodynamic and aeroacoustic analyses based on boundary element method solutions. In order to capture the aerodynamic influence of MiTEs through potential-flow simulations, the MiTEs are replaced by trailing edge plain flaps which provide equivalent aerodynamic responses. Results concerning the proposed controller capability to alleviate high-frequency blade loads and subsequent emitted noise from BVI events are presented and discussed.

Validations of Rotor Load Analysis Using Flexible Multibody Dynamics with Multiple Wake Panels for Low-Speed Flights

Rotor load predictions of the UH-60A helicopter in low-speed flights are validated using nonlinear flexible multibody dynamic analyses with multiple wake panels. A nonlinear flexible multibody dynamic analysis code, DYMORE II, is used for predicting the blade section airloads and structural loads of the UH-60A rotor in low-speed forward and descending flights at the advance ratio of 0.15. The freewake model with multiple wake panels is integrated implicitly into DYMORE II to represent effectively the BVI (blade–vortex interaction) phenomenon of the rotor in low-speed flights. For both low-speed forward and descending flights, the multiple wake panel and single wake panel models both validate the blade section normal forces (or lift forces) moderately or well with the flight test data; however, they predict the blade section pitching moments moderately or poorly. For blade structural loads, two analyses using the multiple wake panel and single wake panel models validate the oscillatory flap bending moments moderately for both flight conditions; however, they correlate the oscillatory lead–lag bending moments poorly with the measured data. In addition, the multiple wake panel and single wake panel models both validate the oscillatory torsion moment well for the low-speed descending flight; but they predict it moderately or poorly for the low-speed forward flight. Through the present study, the analyses using multiple wake panels show better predictions of the rotor loads in low-speed flights as compared to the results with a single wake panel, along with efficient and reasonable computation resource and time.

The second wind-tunnel test of DLR’s multiple swashplate system

After proving its individual blade control (IBC) capabilities on a four-bladed rotor system in the wind tunnel, the DLR’s multiple swashplate system and its control software were developed further to allow IBC operation on a five-bladed rotor system. After preliminary tests in hover conditions, a second wind-tunnel test was performed in the DNW-LLF wind tunnel in late 2016. The goal of this test on a Mach-scaled model rotor was to reduce vibration, noise, and required rotor power in different flight conditions using proven IBC strategies as well as localized pitch control (LPC) on five blades. In descent flight condition, significant reductions of blade–vortex interaction noise relative to the baseline case were achieved on both sides of the rotor disk through the application of 2/rev higher harmonic control as well as using an LPC schedule. Through the application of 3/rev IBC as well as a vibration controller using a 4–6/rev multi-harmonic IBC signal, 5/rev hub loads were reduced significantly and the 5/rev vertical vibrations were nearly eliminated. In addition, during simulated high-speed-level flight with a wind speed of 76 m/s, the required rotor power was successfully reduced using a 2/rev input with an amplitude of 1 $$^\circ$$ .

Rotorcraft comprehensive code assessment for blade–vortex interaction conditions

The scope of this paper is the presentation of the computational methodologies applied in the comprehensive code for rotorcraft developed in the last years at Roma Tre University, along with the assessment of its prediction capabilities focused on flight conditions characterized by strong blade–vortex interactions. Boundary element method approaches are applied for both potential aerodynamics and aeroacoustics solutions, whereas a harmonic-balance/modal approach is used to integrate the rotor aeroelastic equations. The validation campaign of the comprehensive code has been carried out against the well-known HART II database, which is the outcome of a joint multi-national effort aimed at performing wind tunnel measurements of loads, blade deflection, wake shape and noise concerning a four-bladed model rotor in low-speed descent flight. Comparisons with numerical simulations available in the literature for the same test cases are also presented. It is shown that, with limited computational cost, the results provided by the Roma Tre aero-acousto-elastic solver are in good agreement with the experimental data, with a level of accuracy that is in line with the state-of-the-art predictions. The influence of the vortex core modeling on aerodynamic predictions and the influence of the inclusion of the fuselage shielding effect on aeroacoustic predictions are discussed.

Aeroacoustics of a rotor ingesting a planar boundary layer at high thrust

Aeroacoustic measurements and analysis have been made for an unshrouded rotor partially immersed in a planar equilibrium turbulent boundary layer at low Mach number. This configuration provides an idealized model of inflow distortion effects seen when a rotor is mounted adjacent to the hull or fuselage of a vehicle. At low and moderate thrust conditions, the rotor produces broadband noise organized into haystacks produced by large eddies of the ingested turbulence being cut multiple times by successive rotor blades. At high thrust, however, the acoustic signature changes and becomes louder and more tonal. This change is accompanied by separation of the boundary layer from the wall in the vicinity of the rotor blade disk. The separation region is highly unsteady and populated by intense vortex structures. Acoustic analysis suggests that blade–vortex interactions with these structures are the source of the additional tonal noise at high thrust.

Coupling of an Unsteady Aerodynamics Model with a Computational Fluid Dynamics Solver

Momentum source methods are an efficient means of representing airfoils in Navier–Stokes computational fluid dynamics (CFD) simulations. Momentum source terms are added to the Navier–Stokes equations instead of resolving the solid boundary of the airfoil with a mesh. These source terms are calculated using an aerodynamics model. This approach is useful where the overall performance and mid- to far-field influence of a wing or rotor are desired and details of the flowfield near the blade are not the objective of the simulation. One example is simulation of rotorcraft operations where the objective may be to assess an operation’s feasibility in terms of control margins, rather than to inform rotor design decisions. Coupling an aerodynamics model to a CFD solver is straightforward in cases where the airflow relative to the blade is steady. Unsteady conditions require an unsteady aerodynamics model, complicating the coupling with the CFD solver. A coupling method is proposed whereby the incident velocity is extracted from the CFD solution and corrected using a theory-based approximation for the unsteady induced velocity. The steady-state momentum source method is demonstrated for two- and three-dimensional simulations, and the unsteady coupling method is validated against experiments on a pitching airfoil and verified for blade–vortex interactions. The unsteady coupling method enables meaningful incident velocities to be extracted from unsteady flowfields, as shown by agreement with experiments and simulations using analytical expressions for the incident velocity in place of the CFD solver.

Effects of Crosswind Flight on Rotor Harmonic Noise Radiation

To develop recommendations for procedures for helicopter source noise characterization, the effects of crosswinds on main rotor harmonic noise radiation are assessed using a semi-empirical model of the Bell 430 helicopter. Crosswinds are found to have a significant effect on blade-vortex interaction (BVI) noise radiation when the helicopter is trimmed with the fuselage oriented along the inertial flight path, resulting in a “slipped” flight condition. However, the magnitude of BVI noise remains unchanged when the pilot orients the fuselage along the aerodynamic velocity vector, “crabbing” for zero aerodynamic sideslip angle. The effects of wind gradients on BVI noise are also investigated and found to be smaller in the crosswind direction than in the headwind direction. The effects of crosswinds on lower harmonic noise sources at higher flight speeds are also assessed. In all cases, the directivity of radiated noise is somewhat changed by the crosswind. The model predictions agree well with a limited set of flight test data for the Bell 430 helicopter for cruise and approach flight conditions, both measured with and without crosswinds present. The results of this investigation would suggest that flight paths for future acoustic flight testing are best aligned across the prevailing wind direction to minimize the effects of winds on noise measurements when wind cannot otherwise be avoided.

Computational investigation of noise interaction for a nano counter-rotating rotor in a static condition

The interaction between the upper and lower rotors greatly influences the aeroacoustic characteristics of a counter-rotating nano-coaxial rotor. To study this influence, a numerical investigation was carried out. The unsteady flow field of a single upper rotor was first studied with a large-eddy simulation computational fluid dynamics method coupled with a sliding-mesh technique. The Ffowcs Williams–Hawking equation method was used to investigate the aeroacoustic characteristics of the upper rotor based on the flow field. An experimental setup was established to validate the computational approach. The experimental results matched well with the computational results. Additionally, results show that the peak value of the total sound pressure level appeared near the blade tip, which verified that the tip vortex was one of the most important sources of rotor noise. Then the aeroacoustic noise of the nano-coaxial rotor was studied numerically. It was found that the total sound pressure level of the nano-coaxial rotor was greater than that of the upper rotor. Flow field analysis showed that the shedding vortices of the upper rotor interacted with the lower rotor, resulting in a blade–vortex interaction. It was evident that the aeroacoustic noise was enhanced by the interference between the upper and lower rotors.

State-space coaxial rotors inflow modelling derived from high-fidelity aerodynamic simulations

Wake inflow modelling is a crucial issue in the development of efficient and reliable computational tools for flight dynamics and aeroelastic analyses of rotorcraft. The aim of this work is the development of state-space, wake inflow modelling techniques for coaxial rotors perturbed from steady flight conditions, based on simulations provided by high-fidelity aerodynamic solvers. These provide relationships of the coefficients of an approximated linear distribution of wake inflow over upper and lower rotor discs either with rotor controls and helicopter kinematic variables or with thrust and in-plane moments generated by the rotors. A three-step identification procedure is proposed. It consists in: (i) evaluation of wake inflow due to harmonic perturbations of rotor kinematics; (ii) determination of the corresponding inflow coefficients (and rotors loads) transfer functions, followed by (iii) their rational approximation and transformation into time domain to obtain the final differential form. Aerodynamic responses are predicted through a boundary element method tool for potential flows, capable of capturing effects due to wake distortion, multi-body interference (like that in coaxial rotors) and severe blade–vortex interactions. The derived state-space models are validated by correlation with inflow distributions directly calculated by the aerodynamic solver, for a coaxial rotor subject to arbitrary perturbations.

Pressure feedback-based blade–vortex interaction noise controller for helicopter rotors

With the aim of alleviating the noise annoyance emitted by blade–vortex interactions occurring on helicopter main rotors, the present work presents a methodology suitable for the identification of a multi-cyclic harmonic controller based on the actuation of rotor blades equipped with Miniature Trailing Edge Effectors. The objective of the control methodology is the direct suppression of the aerodynamic noise sources by generation of localized high-harmonic blade–vortex interaction counter-actions. The set-up of control devices is selected on the basis of the blade–vortex interaction scenario, taking into account a trade-off between effectiveness and power requirement. The control law is efficiently identified by means of an optimal controller synthesized through suitable two-dimensional multi-vortex, parallel blade–vortex interaction problems. The proposed methodology is validated by the application to realistic helicopter main rotors during low-speed descent flights, numerically simulated through high-fidelity aerodynamic and aeroacoustic solvers based, respectively, upon a three-dimensional free-wake boundary element method to solve the potential flow around rotors in blade–vortex interaction conditions and the Farassat 1A formulation. Results concerning the capability of the proposed controller to alleviate the blade–vortex interaction noise emitted by a realistic helicopter main rotor are presented and discussed.

Development and validation of a hybrid method for predicting helicopter rotor impulsive noise

Prediction of helicopter rotor impulsive noise is practically a very challenging task. This paper describes a hybrid method to predict rotor impulsive noise for both high-speed impulsive noise and blade–vortex interaction noise. The hybrid solver has been developed by combining the advantages of three different methods: (1) a computational fluid dynamics method based on Reynolds-averaged Navier–Stokes equations to account for the viscous and compressible effects near the blade; (2) a vorticity transport model to predict rotor wake system without artificial dissipation; and (3) an acoustic calculation method, based on Ffowcs-Williams Hawkings equation implemented to a permeable data surface. The developed hybrid solver is validated through available test data, for the cases of UH-1H model rotor, AH-1 Operational Loads Survey rotor, and Helishape 7A rotor. Peak sound pressure level of high-speed impulsive noise is accurately predicted with relative errors less than 7%. Additionally, acoustic waveform of blade–vortex interaction noise is well captured though it is sensitive to small changes in aerodynamic load. It is suggested that present hybrid method is versatile for the prediction of rotor impulsive noise with moderate computational cost.

The effects of dynamic-stall and parallel BVI on cycloidal rotor

PurposeThis paper aims to investigate the mechanisms lying behind the cycloidal rotor under hovering status.Design/methodology/approachExperiments were conducted to validate the numerical simulation results. The simulations were based on unsteady Reynolds-averaged Navier–Stokes (URANS) equations solver and the sliding mesh technique was used to model the blade motion. 2D and 2.5D simulations were made to investigate the 3D effects of turbulence. The effects of pressure and viscosity were compared to study the significance of the blade motion on force generation.FindingsThe 2.5D numerical simulation cannot produce more accurate results than the 2D counterpart. The pitching motion of the blade results in dynamic stall. The dynamic stall vortices induce parallel blade vortex interaction (BVI) upon downstream blades. The interactions between the blades delay the stall of the blade which is beneficial to the thrust generation. The blade pitching motion is the dominant contributor to the force generation and the turbulence is the secondary. Strong downwash in the rotor cage varied the inflow velocity as well as the effective angle of attack (AOA) of the blade.Practical implicationsCycloidal rotor is a propulsion device that can provide omni-directional vectored thrust with high efficiency and low noise. To understand the mechanisms lying behind the cycloidal rotor helps the authors to design efficient cycloidal rotors for aircraft.Originality/valueThe authors discovered that the blade pitching motion plays primary role in force generation. The effects of the dynamic stall and BVI were studied. The reason why cycloidal rotor can be more efficient was discussed.

A Vorticity-preserving Hydrodynamical Scheme for Modeling Accretion Disk Flows

Vortices, turbulence, and unsteady nonlaminar flows are likely both prominent and dynamically important features of astrophysical disks. Such strongly nonlinear phenomena are often difficult, however, to simulate accurately, and are generally amenable to analytic treatment only in idealized form. In this paper, we explore the evolution of compressible two-dimensional flows using an implicit dual-time hydrodynamical scheme that strictly conserves vorticity (if applied to simulate inviscid flows for which Kelvin's Circulation Theorem is applicable). The algorithm is based on the work of Lerat et al., who proposed it in the context of terrestrial applications such as the blade–vortex interactions generated by helicopter rotors. We present several tests of Lerat et al.'s vorticity-preserving approach, which we have implemented to second-order accuracy, providing side-by-side comparisons with other algorithms that are frequently used in protostellar disk simulations. The comparison codes include one based on explicit, second-order van Leer advection, one based on spectral methods, and another that implements a higher-order Godunov solver. Our results suggest that the Lerat et al. algorithm will be useful for simulations of astrophysical environments in which vortices play a dynamical role, and where strong shocks are not expected.

From Aeroacoustics Basic Research to a Modern Low-Noise Rotor Blade

In 2015, Airbus Helicopters unveiled the secrecy around its Dauphin successor and presented the all new H160 helicopter. A special feature that immediately attracted attention was its unusual and revolutionary fore-aft swept main rotor blade. This design, which aims to significantly reduce the blade-vortex interaction noise signature and reduce high-speed forward flight power requirements, was first investigated in the 1990s. At this time, DLR and ONERA formed a joint team to acoustically optimize a rectangular reference rotor blade. Based on state-of-the-art comprehensive rotor codes and a 50/50 work share, the ERATO rotor blade design was developed, patented worldwide and tested on both a rotor test rig and in the wind tunnel. Airbus Helicopters (then: Eurocopter) adopted the design, optimized hover and forward flight high lift performance and prepared it for serial production, as the Blue EdgeTM rotor blade on the Airbus Helicopters H160 helicopter.

A 3D analytical model for orthogonal blade-vortex interaction noise

A 3D analytical model of an Orthogonal Blade-Vortex Interaction (OBVI) for Counter-Rotating Open Rotor (CROR) tonal noise is investigated. The specific influence of two parameters taking into account the three-dimensionality of both the vortex velocity and the convection velocity within the rotor-rotor volume is addressed. The first step is to extract the vortex parameters from a recent unsteady Reynolds-Averaged Navier-Stokes computation and validate different vortex models. Lamb-Oseen and Scully vortices reproduce the behavior of the tip-vortex tangential velocity fairly well. Regarding the vortex axial velocity modeling, a Gaussian profile fits well with numerical results. On the one hand, the impact of the stream-tube contraction unbalances the lobes of the unsteady pressure with opposite phases produced by the OBVI event. This effect is larger than that of an equivalent blade sweep. On the other hand, adding the axial velocity deficit to the tangential one also unbalances the pressure lobes. Finally, from an acoustic point of view using Curle's acoustic analogy, both the stream-tube contraction and the axial velocity deficit have the same effect: they turn an acoustically-low efficient quadrupole into a strong dipole making these parameters fundamental for future CROR OBVI investigations.