Rotor Blade-vortex Interaction Research Articles

Numerical Investigation of an Unsteady Blade Surface Blowing Method to Reduce Rotor Blade-Vortex Interaction Noise

This paper presents a numerical investigation of unsteady surface blowing using periodic variations of jet velocity with azimuthal angle to reduce helicopter rotor blade-vortex interaction (BVI) noise. The unsteady blowing is modeled as the mass flow outlet boundary condition of time-varying jet velocity on the blade surface grid using computational fluid dynamics. The same high-resolution overset grid system and flow/noise solver are used to perform a detailed flow field simulation and noise prediction for the nonblowing baseline case and the steady/unsteady blowing cases under the rotor BVI condition, and a grid convergence study for the steady and unsteady blowing cases is carried out. The BVI noise reduction and rotor thrust coefficient results of the unsteady blowing method and the previously published steady blowing with constant jet velocity are then compared. The noise reduction level of unsteady blowing is approximately equivalent to that of steady blowing (noise reduction is more than 3 dB). However, the loss in rotor thrust coefficient caused by unsteady blowing (3.3%) is only half of that by steady blowing (6.3%); the air mass cost by unsteady blowing is only 63.7% of that by steady blowing per rotation revolution. The results show that unsteady blowing can effectively reduce BVI noise with lower cost and less thrust loss.

Application of High-Order WENO Scheme in the CFD/FW–H Method to Predict Helicopter Rotor Blade–Vortex Interaction Tonal Noise

The accurate prediction of helicopter rotor blade–vortex interaction (BVI) noise is challenging. This paper presents an implementation of the seventh-order improved weighted essentially non-oscillatory (WENO-Z) scheme for predicting rotor BVI noise using a high-resolution numerical method based on the Reynolds-averaged Navier–Stokes and the Ffowcs Williams–Hawkings equations. The seventh-order improved WENO-Z scheme is utilized to minimize the inherent numerical dissipation of the reconstruction method in the monotone upstream-centered scheme for conservation laws (MUSCL), thereby improving the rotor wake resolution and the BVI noise-prediction accuracy. The three-layer dummy cell method is used to ensure that the flux at the boundary maintains seventh-order accuracy. The effectiveness of the flow solver and the acoustic solver is validated using the Helishape-7A rotor and the UH-1H rotor, respectively. The flow field and BVI noise characteristics of the OLS rotor obtained from the fifth- and seventh-order WENO-Z schemes are compared with that of the third-order MUSCL for coarse and fine background grids. The wake resolution, noise-prediction accuracy, and computational cost of the three schemes are compared. The results show that the high-order WENO scheme provides higher accuracy for flow field simulation and BVI noise prediction than the MUSCL, but the computational cost of the WENO scheme increases substantially as the grid resolution increases. However, the WENO scheme can predict BVI using a coarser grid than the MUSCL. The computational cost of the WENO scheme is relatively low under the same flow field simulation resolution.

Parametric Effect of Blade Surface Blowing on the Reduction of Rotor Blade–Vortex Interaction Noise

AbstractReducing rotor blade–vortex interaction (BVI) noise is a major challenge in rotorcraft research. This paper presents a high-resolution numerical study of the effects of blade surface blowin...

Numerical investigation on noise reduction of rotor blade-vortex interaction using blade surface jet blowing

The effects of blade surface jet blowing on the reduction of rotor blade-vortex interaction (BVI) noise are investigated using the compressible Reynolds-averaged Navier-Stokes equations and the Ffowcs Williams-Hawkings equations. A detailed analysis of the accuracy of the Weighted Essentially Non-Oscillatory (WENO) scheme and Monotone Upstream-Centered Scheme for Conservation Laws for simulating the flow field and predicting the BVI noise is performed using a two-bladed scaled AH-1 helicopter main rotor undergoing BVI as the baseline case. Additionally, a grid convergence study is conducted. The BVI noise is accurately predicted using the fifth-order WENO scheme and high-resolution grid system (the prediction errors of the peak sound pressure are less than 4%). Surface boundary conditions are introduced to model the effect of jet blowing on the blade surface on the flow control of the baseline case. Blade surface blowing is analyzed near the trailing edge and the leading edge near the rotor tip. The results show that blowing near the blade trailing edge is effective for BVI mitigation by altering the tip vortex strength and increasing the blade-vortex miss distance. The sound pressure level is reduced by 2.6 dB, and the peak amplitude is attenuated by more than 30%, with a 9.8% loss in rotor thrust. Therefore, blowing near the trailing edge may be considered a promising approach for active BVI noise control.

Helicopter Blade-Vortex Interaction Airload and Noise Prediction Using Coupling CFD/VWM Method

As a high resolution airload with accurate rotor wake is pivotal for rotor BVI (Blade-vortex interaction) analysis, a hybrid method with combined Navier-Stokes equation, viscous wake model, and FW-H (Ffowcs Williams-Hawkings) equation is developed for BVI airload and noise in this paper. A comparison with the CFD (Computational Fluid Dynamics)/FW-H method for the AH-1/OLS (Operational Load Survey) rotor demonstrates its capability for favorable accuracy and high computation efficiency. This paper further discusses the mechanisms for the impacts of four flight parameters (i.e., tip-path-plane angle, thrust coefficient, tip Mach number, advance ratio) on BVI noise. Under the BVI condition, several BVI events concurrently occur on the rotor disk. Each interaction has a distinct radiation direction which depends on the interaction azimuth, and its noise intensity is highly associated with the characteristic parameters (e.g., miss-distance, interaction angle, vortex strength). The BVI noise is dominated by the interactions at 30–90° in azimuth on the advancing side, of which the wake angle range is from 180° to 540°. Furthermore, the tip-path-plane angle, thrust coefficient, and tip Mach number change the noise intensity mainly via miss-distance, interaction angle, and vortex strength, but for different advance ratios, the noise intensity and propagation direction are more dependent on the interaction angle and interaction azimuth.

Synthesis of Active Twist Controller for Rotor Blade–Vortex Interaction Noise Alleviation

This paper deals with an efficient computational process for the synthesis of a low-frequency controller aimed at alleviating helicopter rotor blade–vortex interaction noise. It is based on an optimal, multicyclic, control approach that exploits distributed torque loads actuated by smart materials to twist rotor blades at the 2/rev frequency (active twist rotor concept). Aeroacoustic rotor simulations are based on aerodynamic and aeroelastic tools capable of accurately predicting wake–blade miss distance, which plays a crucial role in blade–vortex interaction noise emission. Its modification is the main objective of the proposed controller action. The control law is identified through a numerically efficient aeroacoustic solver based on analytical–numerical sectional aerodynamics modeling. Numerical investigations first examine noise emission sensitivity to active twist actuation, then identify control and output variables suitable for the closed-loop controller. Finally effectiveness of the proposed low-frequency feedback control law for blade–vortex interaction noise alleviation is assessed by application to helicopter rotor in descent flight.

An Experimental and Theoretical Study of Blade–Vortex Interaction Noise

An experimental method to study helicopter blade–vortex interaction (BVI) noise has been developed. Called the blade-controlled disturbance interaction, the method nearly replicates a rotor BVI by having a single-bladed rotor pass through a stationary gust field specially designed to simulate the highly impulsive induced velocity field of a vortex. This approach can be used to evaluate the effectiveness of blade design changes to noise radiation during a BVI event, as well as to study the effect of different interaction geometries on the resulting noise. The first set of experiments in this facility shows that the directionality of BVI noise radiation is very sensitive to the interaction angle. Oblique interactions spread the acoustics energy over wider azimuth angles compared to parallel interactions, as well as move the peak noise location closer to the rotor plane. Linear two-dimensional unsteady aerodynamic theory predicts the overall directionality trends reasonably well. However, the actual pulse shapes do not match and noise levels are underpredicted, particularly for oblique interactions.

Time-Domain Approach for Acoustic Scattering of Rotorcraft Noise

This paper addresses acoustic scattering of rotorcraft noise in the time domain. A time-domain equivalent source method is used since it is considered to be a computationally efficient method to solve acoustic scattering. In addition, the time-domain method provides a solution for all frequencies of interest in a single computation and is able to predict the acoustic scattering of aperiodic signals. The prediction is validated against exact solutions for a monopole source. The numerical method is then used to predict acoustic scattering of noise from a BO105 tail rotor by a representative fuselage. Complex directivity patterns are seen in the near field, and a large scattering effect is observed in the far field to the side of the fuselage. The time-domain code results of sound pressure level are validated against the results obtained by a frequency-domain analysis. Finally, acoustic scattering for an impulsive noise source is investigated to simulate main rotor blade–vortex interaction noise. The scattered pressure has a comparable amplitude as that of the incident pressure so that the total pressure is dramatically changed compared to the incident pressure. For the impulsive noise, a large computational time saving is achieved compared to the frequency-domain approach, in which the computation must be repeated for each frequency.

A Parameter Identification Method for Helicopter Noise Source Identification and Physics-Based Semiempirical Modeling

A new physics-based parameter identification method for rotor harmonic noise sources is developed using an acoustic inverse simulation technique. This new method allows for the identification of individual rotor harmonic noise sources and allows them to be characterized in terms of their individual non-dimensional governing parameters. This new method is applied to both wind tunnel measurements and ground noise measurements of two-bladed rotors. The method is shown to match the parametric trends of main rotor Blade-Vortex Interaction (BVI) noise, allowing accurate estimates of BVI noise to be made for operating conditions based on a small number of measurements taken at different operating conditions.

An Analytical Study of Parametric Effects on Rotor—Vortex Interaction Noise

A numerical method for modelling the aerodynamic interaction between rotor blades and an independently generated vortex is developed. In this method, the Navier—Stokes/Euler equations are adopted for capturing the features of rotor blade—vortex interaction (BVI) flow and the overset grids technique is used for multi-zone solution. To incorporate the effects of the vortex in the solution, a generalized grid-velocity approach is presented to simplify the computations. The acoustic prediction is based on the well-known Ffowcs Williams-Hawkings (FW-H) equation for acoustic pressure. By this method developed, a parametric study is performed to examine the effects of the vortex strength, miss-distance, core-radius, oblique interaction, and interaction angle in a blade-shaft plane on BVI noise. The results indicate that the increase in miss-distance, core-radius, obliqueness of interaction, or interaction angle in a blade-shaft plane is an effective approach to reduce BVI noise, makes the BVI noise less impulsive (lower high harmonic content in noise energy) and increases the high noise encompassed area. However, the decrease in vortex strength changes only the magnitude of noise, not its distribution in different harmonics. The research also indicates that oblique interaction in a rotor plane has a significant effect on radiation directionality, while other parameters show little impact on it.

Computational Modeling of HART II Blade-Vortex Interaction Loading and Wake System

Correlations using a loosely coupled trim methodology of the computational fluid dynamics (OVERFLOW-2) and computational structural dynamics (CAMRAD-II) codes are presented to calculate the helicopter rotor blade-vortex interaction airloads and wake system for the higher-harmonic aeroacoustic rotor test (HART II) rotor at an advance ratio of 0.15. Five different grid models are studied to quantify the effects of grid refinement on rotor-wake resolution. The fine grid model has a total of 113 million grid points and it improves airload predictions compared with the standard grid model for three HART II test cases: baseline, minimum noise, and minimum vibration. The rotor-wake positions are well predicted by this fine grid model. The computed vorticity field for a young vortex using the fine grid model is compared with the measured particle image velocimetry data and the results are good. The fine grid model underpredicts the experimental value for the maximum vorticity by 61%. The predicted vortex core radius is Is % in chord for the fine grid while the measured data show about 5% chord length. The predicted swirl velocity is, however, higher than the measured data for this vortex. The results in this paper provide the first quantitative comparisons between the measured and computational fluid dynamics/computational structural dynamics computed flowfield for a helicopter rotor-wake system.

Analysis of SPR measurements from HART II

The wind tunnel test of HART II (Higher Harmonic Control Aeroacoustic Rotor Test), performed in October 2001 in the Large Low-Speed Facility (LLF) of the German-Dutch Wind-tunnel (DNW), is part of an international cooperative program by the German DLR, French ONERA, DNW, NASA Langley and the US Army Aeroflightdynamics Directorate (AFDD). The main objective of the program is the investigation of rotor wake and its influence on rotor blade–vortex interaction (BVI) noise with and without higher harmonic pitch control (HHC). For blade position and deflection measurements the Stereo Pattern Recognition (SPR) technique was used for the first time. This technique is based on a 3-dimensional reconstruction of visible marker locations by using stereo camera images. An evaluation of these images leads to the spatial position of markers which are attached to each of the four blades and to the bottom of the fuselage and thus to the blade motion parameters in flap, lead-lag and torsion. In this paper the different analysis methods and post-processing of SPR data and the final results for all four rotor blades are presented and the advantages, drawbacks and the potential of this technology are shown.

A wind-tunnel based study of helicopter tail rotor blade vortex interaction

AbstractThe interaction of a helicopter tail rotor blade with the tip vortex system from the main rotor is a significant source of noise and, in some flight states, can produce marked reductions in control effectiveness. This paper describes a series of wind-tunnel tests to simulate tail rotor blade vortex interaction with a view to providing data for the development and validation of numerical simulations of the phenomenon. In the experiments, which were carried out in the Argyll wind-tunnel of Glasgow University, a single-bladed rotor located in the tunnel’s contraction was used to generate the tip vortex which travelled downstream into the working section where it interacted with a model tail rotor. The tail rotor was instrumented with miniature pressure transducers that measured the aerodynamic response during the interaction. The results suggest that the rotor blade vortex interaction is similar in form to that measured at much higher spatial resolution on a fixed, non-rotating blade. The combination of the two datasets, therefore, provides a valuable resource for the development and validation of predictive schemes.

Surface Pressure Measurements of the Orthogonal Vortex Interaction

A unique vortex generator that simulates the trailing vortex produced by a helicopter main rotor in forward flight was used in a study of main rotor/tail rotor blade vortex interaction. Velocity measurements were performed using hot wire anemometry and particle image velocimetry, which showed the interacting vortices to be three dimensional and to contain a significant core axial flow. Surface pressure measurements during vortex collision with an instrumented, stationary blade were obtained at a variety of incidence settings. At the leading edge, a suction peak occurred where the core flow was directed away from the surface and, similarly, a pressure peak occurred where the core flow was directed toward the surface. This behavior can be explained in terms of the impulsively blocked vortex core axial flow. Also, the leading edge suction and pressure peaks were amplified or attenuated by changing the incidence of the blade. However, the changes in surface pressure distribution were such that the peak normal force was approximately the same for each incidence setting

Effects of Surface Blowing/Suction on the Aerodynamics of Helicopter Rotor Blade-Vortex Interactions (BVI) — A Numerical Simulation

Effects of Surface Blowing/Suction on the Aerodynamics of Helicopter Rotor Blade-Vortex Interactions (BVI) — A Numerical Simulation

Experimental investigation of the perpendicular rotor blade-vortex interaction at transonic speeds

Transonic perpendicular rotor blade-vortex interaction (BVI) tests at Mach numbers ranging from 0.68 to 0.9 and Reynolds numbers (based on the airfoil chord) of 3.8-5.5 million were conducted in the UTA highReynolds number, transonic Ludwieg-tube wind tunnel. The scheme involved positioning a lifting wing (vortex generator) upstream of an instrumented NACA 0012 airfoil so that the trailing vortex interacted with the downstream airfoil. Tests were performed at several vortex strengths as well as several vortex core heights above the downstream airfoil. The results obtained from these experiments indicate that a substantial change in the pressure distribution of the downstream airfoil occurs, but most of the effects were confined to the leading 30% of the airfoil chord. A spanwise drift of the vortex core as it passes over the trailing airfoil, similar to results observed previously in low-speed wind tunnel tests, as well as a high degree of unsteadiness in the vicinity of the vortex center were observed.