Computational Evaluation of Controlling Flap-Edge Vortex and Tip Vortex Using Circulation Control Techniques

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.

Rotor blade–vortex interaction noise

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 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.

Computation of circulation control airfoil flows

The geometric effects of Coanda trailing edges on the aerodynamic performance of an airfoil are numerically evaluated for a range of different freestream Mach numbers and momentum coefficients. A Circulation control (CC) airfoil with a circular trailing edge (ACTE) proves to have better control effectiveness at low subsonic freestream speeds (Mach = 0.1). A CC airfoil having an elliptic trailing edge (AETE) outperforms the ACTE at high subsonic flow conditions. The occurrence of C μ-stall for the AETE is greatly postponed, and meanwhile the maximum net lift coefficient increment achieved for the AETE (Δ C L = 0.51) is slightly higher than that of the ACTE (Δ C L = 0.50) at Mach 0.6. Compared to the ACTE, the AETE is found to have better control consistency at different operating velocities and better control stability when the Coanda jet is supersonic. Through careful consideration of the aerodynamic performance and the control effects, the most appropriate axial ratio for an AETE ellipse is within the interval from 1.5 to 2. Finally, the flow field instability phenomenon and the jet detachment induced by the supersonic Coanda jet are investigated. A self-sustained shock-wave instability phenomenon without jet detachment is first observed in this paper.

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.

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.

[Abstract] An experimental study has been conducted to examine the effects that Gurney flap span has on the lift and drag of the wing-flap configuration and the near-field trailing vortex strength. The configurations tested included a rectangular wing, a full-span Gurney flap and two Gurney flaps having spans equal to one half the wing span, one positioned outboard, and the other inboard. All three flaps had a height equaling six percent of the wing chord. As expected, all three flaps significantly changed the wing lift and drag characteristics. While the flapped configurations developed much higher both lift and drag coefficients, the overall effect on the wing lift-to-drag ratio has been found to be detrimental for all three flap configurations as compared to the baseline wing, the negative effect increasing with the flap span. Also, significant flattening of the lift-to-drag vs. lift coefficient curve has been observed. As expected, the addition of the flaps led to additional trailing vortices which affect the overall trailing vortex rollup and strength. The effect of the flaps on the trailing vortex strength in the near field has been quantified using a rollup tightness factor. This measure has indicated that while the configurations employing the full-span Gurney flap and the 50-percent-span flap located over the outboard wing sections exhibited trailing vortex strength reductions over most of the range of angles of attack tested, the configuration involving the inboard 50-percent flap developed stronger trailing vortices in the near field. These results have indicated that an effective means for trailing vortex reduction via Gurney flap use should be sought within the design space incorporating flaps located outboard. Future studies should include outboard Gurney flaps of varying relative span, other wing planforms and higher Reynolds numbers. Also, far field effects need to be examined.

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.

Helicopter Blade-Vortex Interaction (BVI) generally occurs under certain conditions of powered descent or during extreme manoeuvring. The vibration and acoustic problems associated with the interaction of rotor tip vortices and the following blades is a major aerodynamic concern for the helicopter community. Numerous experimental and computational studies have been performed over the last two decades in order to gain a better understanding of the physical mechanisms involved in BVI. The most severe interaction, in term of generated noise, happens when the vortex fllament is parallel to the blade, thus afiecting a great portion of it. The majority of the previous numerical studies of parallel BVI fall within a potential ∞ow framework therefore precluding vortex-induced ∞ow separation. Some Navier-Stokes approaches using dissipative numerical methods in conjunction with RANS-type turbulence models have also been attempted, but with limited success. In this work, the situation is improved by increasing the fldelity of both the numerical method and the turbulence model. A kineticenergy conserving flnite-volume scheme using a collocated-mesh arrangement, specially designed for simulation of turbulence in complex geometries, was implemented. For the turbulence model, a costefiective zonal hybrid RANS/LES technique is used. The concentrated tip vortex is not attenuated as it is convected downstream and over a NACA-0012 airfoil. The lift, drag, moment coe‐cients induced by the passage of the vortex are monitored in time and compared with available experimental data.

Circulation control can greatly enhance lift on a xed wing. However, the introduction of a high-speed Coanda jet is theorized to signi cantly increase noise, which is a growing concern for a variety of potential circulation control applications. The sound produced by a circulation control airfoil is measured using a phased microphone array and freeeld microphones in an open-jet anechoic wind tunnel. The phased array maps reveal dominant high-frequency noise sources at the sidewall-trailing edge interfaces. These three-dimensional sources prohibit the accurate measurement of two-dimensional circulation control trailing edge noise by a single microphone. Hence, an integrated phased array spectrum is computed and compared with a theory that identi es three potential noise sources: the interaction of the freestream boundary layer with the slot lip edge, the interaction of boundary layer turbulence with the rounded trailing edge, and the interaction of the jet with the surrounding slot geometry. The theory requires detailed ow eld information, including friction velocities, boundary layer thicknesses, displacement thicknesses, and local mean velocities. These parameters are estimated using particle image velocimetry measurements. A quantitative comparison of the measured and theoretical spectra reveal large di erences. The model predicts that the interaction of the freestream boundary layer with the slot lip is the dominant noise source mechanism over most frequencies.

: 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.

The flows about 2-D and 3-D bluff trailing edge circulation control (CC) airfoils are computed using steady Reynolds Averaged Navier-Stokes (RANS) methods. The 2-D foil is the NCCR 1510-7067 elliptical CC airfoil with circular and logarithmic spiral trailing edge geometries. The free stream Reynolds number, based on chord, is 5.45 × 10 5 , with a free stream Mach number of 0.12. For the circular trailing edge the slot height, blowing rate and angle of attack are varied, while for the logarithmic spiral only the blowing rate is varied. The 3-D foil is a semi-span wing with an elliptical cross section. It is run with a chord-based Reynolds number of 2 × 10 6 and two blowing rates. The 2-D flows are computed using the compressible, segregated solver, Fluent. 2-D results show that the full-Reynolds stress turbulence model (FRSM) predicts the correct jet detachment behavior for the circular trailing edge although the integrated lift forces are consistently underpredicted. The coanda jet detachment point for the logarithmic spiral trailing edge is predicted correctly for a lower blowing rate, but as blowing rate increases, the jet does not detach until it has wrapped around to the pressure side. We show additional 2-D results using mesh refinement via grid adaption and isotropic eddy viscosity turbulence models. The 3-D simulations use the incompressible segregated Fluent solver applying the k −ω SST turbulence model. Results show a slight attachment of the the coanda jet on the pressure side, but the results are generally encouraging.