Rotor blade–vortex interaction noise

The growth that aviation has seen in the last decades has drawn the attention on the environmental impact of aircraft. An important part of this environmental impact is the noise emitted by air vehicles, which is considered rather significant for community annoyance. The generation and propagation of air vehicle noise are two different areas of interest, which require accurate prediction in order to control the emitted noise levels. The present thesis employs numerical methods in order to investigate various air vehicle noise propagation effects. It is divided in two parts: the far field and the near field study. Each of these studies is concentrated on the sound propagation mechanisms that are dominant for each case and uses a numerical method that is best suited for it in terms of mechanisms incorporated and cost effectiveness. The far-field study of this thesis focuses on the nonlinear propagation of helicopter rotor noise using the Burgers equation, a well known one direction propagation method. The Burgers equation incorporates geometrical spreading, atmospheric absorption and nonlinear distortion effects. Towards this study, the HELISHAPE descent case experimental database is used. Blade Vortex Interaction (BVI) noise, the dominant noise contributor during descent, is mainly examined. It is shown that advancing side BVI noise is affected by nonlinear distortion, while retreating side BVI noise is not. For some frequency bands the difference between linear and nonlinear calculations can be as high as 7 dB. Based on signal characteristics at source, two quantities are derived. The first quantity (termed polarity) is based on the pressure gradient of the source signal and can be used to determine whether a BVI signal will evolve as an advancing or a retreating side signal. The second quantity (termed weighted rise time) is a measure of the impulsiveness of the BVI signal and can be used to determine at which frequency nonlinear effects start to appear. Finally, polarity and weighted rise time are shown to be applicable in cases of BVI noise generated from different blade tips, as well as, in cases of non-BVI noise. However, employment of the Burgers equation can be time consuming to be included in routine calculations. It also requires knowledge of the initial pressure time signal. The power spectrum alone, which is usually known, is not sufficient. In order to ii overcome these difficulties, three prediction methods are presented that are based on the Burgers equation. These are: i) a numerically generated database, ii) correlation equations and iii) the phase assignment method. Near field propagation of air vehicle noise requires different treatment than far field. The effects which are mainly affecting the propagation are geometrical spreading, convection and refraction effects due to the flow field, as well as reflections and diffraction on the air vehicle surfaces. Towards these objectives, a new low-order flow/acoustics interaction method for the prediction of sound propagation and diffraction in unsteady compressible flow using adaptive 3-D hybrid grids is investigated. The total field is decomposed into the flow field described by the Euler equations, and the acoustics field described by the Nonlinear Perturbation equations. The method is shown capable of predicting monopole sound propagation, while employment of acoustics-guided adapted grid refinement improves the accuracy of capturing the acoustic field. Interaction of sound with solid boundaries is also examined in terms of reflection and diffraction. Sound propagation through an unsteady flow field is examined using static and dynamic flow/acoustics coupling demonstrating the importance of the latter. Proof of concept for the new method is provided by its application to the case of a conventional jet transport airplane, examining the effect of flow field and wing shielding on the near field noise levels. During the aforementioned noise investigation and analysis, results on Blade Wake Interaction (BWI) noise were also reached. Presently, the mechanism of BWI noise generation, as well as the corresponding prediction model, are still under consideration. Helicopter rotor BWI noise is known to be significant during take-off and level flight, while less attention has been given to descent flight conditions, where BVI noise is dominant. Through signal analysis of the HELISHAPE descent case acoustic database, the rotor azimuthal region responsible for BWI noise is localized and the dominance of BVI noise in the BWI frequency region is shown. Coherence analysis of the blade pressure data indicate significant chordwise coherence in the 3 to 4 Struhal number range and absence of acoustic dipoles in the BWI frequency range. The findings of this study support BWI prediction models based on Amiet’s theory and suggest that BWI noise can be ignored for predictions of rotor noise in descent flight conditions.

At the most fundamental level, main rotor loading noise is caused by the harmonically-varying aerodynamic loads (acoustic pressures) exerted by the rotating blades on the air. Rotorcraft main rotor noise is therefore, in principle, a function of rotor control inputs, and thus the forces and moments required to achieve steady, or "trim", flight equilibrium. In certain flight conditions, the ensuing aerodynamic loading on the rotor(s) can result in highly obtrusive harmonic noise. The effect of the propulsive force, or X-force, on Blade-Vortex Interaction (BVI) noise is well documented. This paper presents an acoustics parametric sensitivity analysis of the effect of varying rotor aerodynamic pitch hub trim moments on BVI noise radiated by an S-70 helicopter main rotor. Results show that changing the hub pitching moment for an isolated rotor, trimmed in nominal 80 knot, 6 and 12 deg descent, flight conditions, alters the miss distance between the blades and the vortex in ways that have varied and noticeable effects on the BVI radiated-noise directionality. Peak BVI noise level is however not significantly altered. The application of hub pitching moment allows the attitude of the fuselage to be controlled; for example, to compensate for the uncomfortable change in fuselage pitch attitude introduced by a fuselage-mounted X-force controller.

[Abstract] This paper extends our previous work on the two-dimensional Circulation Control (CC) airfoil studies to three-dimensional applications so that the potential benefits of CC techniques on controlling vortex and reducing noise can be evaluated. Two numerical simulations have been done to identify the effects of CC techniques on controlling the flap edge vortex and the tip vortex. The first is a stream-wise tangential blowing on a wing-flap configuration. It is demonstrated that a gradually varied CC blowing can totally eliminate flap-edge vortex, thus reducing flap-edge noise. The second case involves a span-wise tangential blowing over a rectangular wing with a rounded wing tip. It is found that CC blowing cannot totally eliminate the tip vortex. However, it can control and modify the location of the tip vortex to increase the vertical clearances between the wing and the tip vortex, thus reducing the blade vortex interaction (BVI) and the BVI noise.

A novel passive blade–vortex interaction noise reduction concept is proposed that integrates open-closed resonator cavities within helicopter rotor blades. These cavities attenuate blade–vortex interaction noise through two physical mechanisms. First, they weaken the unsteady surface pressure fluctuations caused by the blade–vortex interactions. Secondly, they disrupt the acoustic focusing process associated with blade–vortex interaction noise. A midfidelity computational framework is used to design and evaluate the effectiveness of the treatment. Two variations of the treatment are considered: one consisting of uniform-depth cavities and the other of variable-depth cavities. The effects of varying the spanwise distribution of both treatment variants are evaluated. The most effective treatment design consists of uniformly distributed, variable-depth resonators, and it is predicted to attenuate noise levels below the rotor by over 8 dB. When both treatments are applied uniformly along the blade span, the primary mechanism of noise reduction is through weakening the unsteady surface pressure fluctuations. However, when the treatments are staggered along specific portions of the blade span, noise is attenuated through a reduction in the focusing of blade–vortex interaction noise. The staggered uniform- and variable-depth resonators yield comparable noise benefits because they disrupt the focusing of blade–vortex interaction noise in a similar manner.

A detailed computational study was conducted to evaluate actively-controlled trailing-edge flaps as a noise reduction device. The helicopter noise problem is first introduced, noting that active control of blade-vortex interaction (BVI) noise is a relatively new field for computational research. Two structural models are considered for the present study; the first is a modal approach suitable for four-bladed hingeless rotors. The second is a finite element approach that is used to consider a five-bladed bearingless rotor. First, open-loop control is implemented with actively-controlled flaps (ACFs) and with conventional root-actuated individual blade control. The effectiveness of both techniques at noise reduction, vibration reduction, and rotor power reduction are considered in BVI descent flight. Subsequently, the BVI noise directivity in the near-field and far-field are examined for baseline and dual ACF noise control cases. Finally, reduction of noise and vibration is successfully demonstrated with a single plain flap configuration on a five-bladed bearingless rotor resembling that of the MD-900.

A discrete implementation of the wavelet transform (WT) was used to analyze blade–vortex interaction (BVI) noise in acoustic helicopter noise signals. A BVI detection algorithm was developed which takes advantage of the prominence of BVI noise in certain subbands. Isolated‐BVI signals were constructed by detecting impulsive BVI events and extracting them from the subbands. Four metrics were tested for suitability as BVI estimators. Two of these, the rms levels of the isolated‐BVI signals and a metric computed from the A‐weighted frequency spectra of the isolated‐BVI signals, were seen to characterize BVI noise severity, but were computationally expensive. The other two metrics were computed from the amplitudes of the BVI events in the subbands. The first, based on the exponential behavior of the amplitudes across subbands, failed to characterize BVI noise severity for certain signals and was highly sensitive to noise. The second, computed from the amplitudes of BVI events in a single subband, was an effective BVI estimator and was relatively insensitive to noise. This estimator is especially convenient because it is easily used as a decision threshold in the detection algorithm, thereby making possible simultaneous BVI detection and characterization.

: A higher harmonic control (HHC) investigation was conducted on a full-scale, isolated XV-15 Rotor in helicopter mode in the NASA Ames 80- by 120-Foot Wind Tunnel to independently control noise, vibration, and trim. The higher harmonic blade root pitch was generated using swashplate oscillations. The radiated blade-vortex interaction (BVI) noise footprint was measured on a plane beneath the rotor with eight microphones mounted on an acoustic traverse. Test results showed that HHC is highly effective in reducing BVI noise, achieving a 12 dB reduction in peak noise level within the noise footprint. Blade pressure feedback was demonstrated to be a viable method for closed-loop noise control. Perturbations in trim parameters and test conditions had small to moderate effects on noise reduction with HHC. Some noise reduction was achieved with no increase in vibratory hub loads. Increases in control loads due to HHC generally limited further noise reduction. The vibration controller achieved about 50 percent reduction in vibratory hub loads with control loads limiting the HHC amplitude. An automatic trim controller was demonstrated to be robust, reaching all target thrust and flapping schedules under all conditions tested.

Free-wake models are routinely used in aeroacoustic analysis of helicopter rotors; however, their semiempiricism is accompanied with uncertainty related to the modeling of physical wake parameters. In some cases, analysts have to resort to empirical adaption of these parameters based on previous experimental evidence. This paper investigates the impact of inherent uncertainty in wake aerodynamic modeling on the robustness of helicopter rotor aeroacoustic analysis. A free-wake aeroelastic rotor model is employed to predict high-resolution unsteady airloads, including blade–vortex interactions. A rotor aeroacoustics model, based on integral solutions of the Ffowcs Williams–Hawkings equation, is utilized to calculate aerodynamic noise in the time domain. The individual analytical models are incorporated into an uncertainty analysis numerical procedure, implemented through nonintrusive Polynomial Chaos expansion. The potential sources of uncertainty in wake tip-vortex core growth modeling are identified and their impact on noise predictions is systematically quantified. When experimental data to adjust the tip-vortex core model are not available the uncertainty in acoustic pressure and noise impact at observers dominated by blade–vortex interaction noise can reach up to 25% and 3.50 dB, respectively. A set of generalized uncertainty maps is derived, for use as modeling guidelines for aeroacoustic analysis in the absence of the robust evidence necessary for calibration of semiempirical vortex core models.

Blade Tip flow and Noise Prediction by Large-Eddy Simulation in Horizontal Axis Wind Turbines

Due to the importance of understanding the aeroacoustics of rotorcraft with continually changing noise sources, this paper presents a new technique for source separation from ground-based acoustic measurements. The source separation process is based on combining a time-domain de-Dopplerization method with the Vold-Kalman (VK) order tracking filter approach. This process can extract rotor harmonic noise even when the sources are continuously changing with time, including impulsive events such as Blade Vortex Interaction (BVI) noise. The advantage of this approach over traditional methods such as harmonic averaging is that the phase and amplitude relationship of acoustic signals is preserved throughout the extraction process. The approach is applied to the measured acoustic data from a Bell 430 helicopter. The measured data were separated into main rotor harmonic, tail rotor harmonic, and broadband residual components. For steady-state conditions, the extracted components could be de-propagated to form acoustic spheres showing the directivity of the separated main and tail rotor components. The source separation process was also applied to a maneuvering flight condition. Each component has different pulse shapes and directivity trends, consistent with aeroacoustic theory.

A linear quadratic Gaussian (LQG) controller for active vibratory loads reduction in helicopters is proposed based on a revisited higher harmonic control (HHC) input by active trailing-edge flaps (ATEFs). Conventional individual blade control (IBC) input is redefined using N − 1/rev interblade phase lead, N/rev collective, and N + 1/rev interblade phase lag signals where 1/rev frequency modulation originates from the multiblade coordinate (MBC) transform. A Mach-scaled flap blade is designed and analyzed by the multibody dynamics analysis DYMORE. A linear time-invariant representation is identified from N/rev envelopes of the input and output responses obtained by DYMORE analysis. A matlab/simulink closed-loop control simulation is designed using the identified state-space realization. The N/rev vibratory loads are reduced up to 52% with flap deflections and the linear control results match well with the nonlinear responses obtained from DYMORE. Furthermore, the multivariable closed-loop stability estimated by the loop transfer functions using disk margin analysis reveals sufficient gain and phase margins.

This paper presents a comparison of multi-fidelity approaches for the HART-II rotor noise prediction using the high-fidelity and multi-disciplinary rotorcraft analysis software suite CREATETM-AV Helios. The acoustic simulations are performed using the noise prediction tools AARON/ANOPP2 and PSU-WOPWOP. Both permeable and impermeable formulations of the Ffowcs Williams and Hawkings equation are considered for noise predictions. The HART-II baseline and 3/rev high harmonic control (HHC) minimum vibration cases are simulated and validated against test measurements. Compared with the free-wake simulation, the highfidelity CFD/CSD loose coupling simulations agree better with the measured data but still under-predict the advancing side airloads. Both impermeable and permeable surface approaches show better noise predictions than the free-wake approach in terms of noise directivity and amplitude. As the finest wake-grid resolution reduces from 7.5% to 5% chord, the impermeable surface strategy shows an increment of 3 dB in maximum mid-frequency sound pressure level (SPL) on the advancing side of the mid-frequency SPL contour. The maximum mid-frequency SPL on the retreating side of the vehicle is only increased by 1 dB. At a finer wake-grid resolution, the permeable surface approach does not show improvement in blade-vortex interaction (BVI) noise prediction on the advancing side of the rotor disk.

An acoustics test of a 2/5 scale model BO-105 helicopter main rotor was conducted in the Duits-Nederlandse Windtunnel (DNW). A range of operating conditions was tested from hover to moderately high flight speeds for various climb and descent rates at different thrust settings. Diagnostic tests including rotor speed and blade geometry changes were made to better isolate and study particular broadband self noise sources. Acoustic data in the form of acoustic pressure time histories and power spectra are used to demonstrate the regions of importance of the different broadband noise sources and their sensitivity to operating conditions. To help interpret the data, comparisons are made to predictions of rotor broadband noise. The predictions are based on self noise data previously obtained from isolated airfoil sections and the use of the NASA ROTONET program to define rotor performance and to sum contributions of noise from individual blade segments. An important result herein is the identification and articulation of a previously unheralded rotor broadband noise source. This source is blade-turbulent wake interaction (BWI) noise which dominates the spectra in the mid-frequencies for off-peak blade-vortex interaction (BVI) noise flight conditions.

This present study investigates the effect of blade tip configurations, such as the sweepback angle and anhedral angle, on the performance and hub vibratory loads for the lift-offset coaxial rotor of a 30,000-pound-class high-speed long-range utility helicopter. The rotorcraft comprehensive analysis code, CAMRAD II, is utilized to conduct the performance and hub vibratory load analyses for the present lift-offset coaxial rotor. The total rotor thrust, torque, and individual rotor’s hub pitch moment and hub roll moment are considered the trim targets. The general properties for the lift-offset coaxial rotor are designed from the X2TD, S-97 Raider, and SB > 1 Defiant, which are lift-offset compound helicopters. The rotor performance and hub vibratory loads are studied with the various blade tip configurations including the sweepback angle and anhedral angle. The rotor power when the rotor blade tip considers only the sweepback angle (20°) is lower than the baseline rotor model by 41.25% at 170 knots. The maximum rotor effective lift-to-drag ratio (L/De) for the lift-offset coaxial rotor using only the sweepback angle and the rotor with both sweepback (20°) and anhedral angles (10°) at 170 knots increase by 10.82% and 5.02%, respectively, compared with the baseline rotor model without both sweepback and anhedral angles. The vibration index (VI) for the rotor with only the sweepback angle is higher than that for the baseline rotor model without both sweepback and anhedral angles by 37.14%. Furthermore, when the rotor blade tip has the anhedral angle, the magnitude of the Blade Vortex Interaction (BVI) decreases compared with the rotor without the sweepback and anhedral angles.

Vortex interaction with a rotating blade is fundamentally important in understanding the unsteady aerodynamics and aeroacoustics problems arising in a helicopter flight. The blade-vortex interaction (BVI) can occur when the main- and/or tail-rotor blades interact with tip vortices previously shed by preceding high speed blades of the main rotor. Most numerical studies have focused on limiting cases where an idealised vortex interacts with a single, non-rotating blade. Despite the significant differences between a real case scenario of BVIs involving more complex characteristics of blade tip vortices and rotating blades, numerical and computational studies nevertheless provide great insight into the deeper understanding of the fluid mechanics that are relevant to the problem. Our computational modelling approaches this problem in several steps. Firstly the ‘impulsive’ instantaneous blocking of the column vortex by a flat plate. This was studied theoretically for incompressible motion byMarshall [1], who produced an elegant model for ‘area change’ waves in the vortex core. Our Computational Fluid Dynamics (CFD) study essentially models this, together with the true compressible pressure wave formation that also arises. In the second step we extend this to the gradual cutting of the vortex by a sharp flat plate that moves, at a finite speed, through the vortex. Finally, we extend this to the gradual cutting of the vortex by a blunt leading edge aerofoil that moves, at a finite speed, through the vortex, which incorporates both the ‘blocking’ and also stretching and distortion of the vortex lines. In this paper, we present only the 3D numerical computations of a single columnar vortex instantaneously ‘blocked’ by a flat plate.