Numerical Simulation of Parallel Airfoil/Vortex Interaction Using a Zonal Hybrid RANS/LES Method

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.

[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.

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.

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.

Rotor blade–vortex interaction noise

Turbulence modelling of low Reynolds number flow effects around a vertical axis turbine at a range of tip-speed ratios

Numerical simulation methods of aerodynamic heating were compared by considering the influence of numerical schemes and turbulence models, and attempting to investigate the applicability of numerical simulation methods on predicting heat flux in engineering applications. For some typical cases provided with detailed experimental data, four spatial schemes and four turbulence models were adopted to calculate surface heat flux. By analyzing and comparing, some influencing regularities of numerical schemes and turbulence models on calculating heat flux had been acquired. It is clear that AUSM+-up scheme with rapid compressibility-modified high Reynolds number k-ω model should be appropriate for calculating heat flux. The numerical methods selected as preference above were applied to calculate the heat flux of a 3-D complex geometry in high speed turbulent flows. The results indicated that numerical simulation can capture the complex flow phenomena and reveal the mechanism of aerodynamic heating. Especially, the numerical result of the heat flux at the stagnation point of the wedge was well in agreement with the prediction of Kemp-Riddel formula, and the surface heat flux distribution was consistent with experiment results, which implied that numerical simulation can be introduced to predict heat flux in engineering applications.

Tip vortex cavitation is the first type of cavitation to take place around most marine propellers. But the numerical prediction of tip vortex cavitation is one of the challenges for propeller wake because of turbulence dissipation during the numerical simulation. Several parameters of computational mesh and numerical algorithm are tested by mean of the predicted length of tip vortex cavtiation to validate a developed method. The predicted length of tip vortex cavtiation is on the increase about 0.4 propeller diameters using the developed numerical method. The predicted length of tip vortex cavtiation by RNG k – e model is about 3 times of that by SST k – ω model. Therefore, based on the validation of the present approach, the cavitating flows generated by two rotating propellers under a non-uniform inflow are calculated further. The distributions of axial velocity, total pressure and vapor volume fraction in the transversal planes across tip vortex region are shown to be useful in analyzing the feature of the cavitating flow. The strongest kernel of tip vortex cavitation is not at the position most close to blade tip but slightly far away from the region. During the growth of tip vortex cavitation extension, it appears short and thick, and then it becomes long and thin. The pressure fluctuations at the positions inside tip vortex region also validates the conclusion. A key finding of the study is that the grids constructed especially for tip vortex flows by using separated computational domain is capable of decreasing the turbulence dissipation and correctly capturing the feature of propeller tip vortex cavitation under uniform and non-uniform inflows. The turbulence model and advanced grids is important to predict tip vortex cavitation.

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.

The impact of 12 spoiler–tab configurations, of different heights and widths, on the tip vortex generated by an oscillating NACA 0015 wing was investigated experimentally. For an oscillating wing equipped with a spoiler, the peak tangential velocity and core and total circulation were greatly reduced compared to a tab, regardless of its width, while the core radius remained largely unaffected with its center displaced vertically above the baseline wing. The most noticeable impact of a spoiler with a reduced height was its potential in alleviating the blade–vortex interaction (BVI) strength. Meanwhile, the largest favorable impact on the critical vortex flow parameters was achieved via a 25%-span spoiler–tab combination with a height of 5 and 2.5% chord, respectively. A contrary effect on the BVI suppression, especially during pitch-up, was, however, observed. The impact on the BVI can be improved by reducing the height of the spoiler at the expense of unfavorable change in the vortex strength and displacement.

The open rotor is being re-considered as an alternative to the advanced turbofan in order to address growing pressure on the aviation industry to reduce fuel consumption and cut carbon dioxide emissions. Meeting future community noise targets is likely to be a challenge for the open rotor, however, and novel noise reduction technology will be required. In this paper, a blade tip winglet on the front rotor is investigated using steady RANS computations for reducing tip vortex interaction noise at takeoff. Only the front rotor row is analysed with the intention being to gain a first insight into the aerodynamics of a blade with a winglet. For this reason, the design of the winglet has not been optimized. Notwithstanding, the winglet design investigated here is predicted to give reductions in maximum tip vortex swirl velocity of order 30% under a takeoff condition representative of modern open rotor operation. The cause of this reduction is shown to be consistent with shedding of bound circulation over a greater spanwise distance. At cruise, the winglet improves the propeller efficiency of the front rotor by 0.4%. A control volume analysis of the lost propeller power shows that approximately half of the propeller efficiency benefit is due to reduced entropy generation associated with the formation of the tip vortices, and a reduction in vortex swirl kinetic energy.

This study explores the movement of an object placed in an air tunnel. In expectation this exploration leads to obtaining a suitable shape of the object and the flow parameters, the object moves along the air tunnel has a particular trajectory. This study used the body of the main object is a cube in which a rectangular wing of aspect ratio 2 is set in the middle of body object. The object can be used as a carrier to transport goods from one place to another through an air tunnel. Three types wing-cube configurations, they are namely wing- cube configuration with airfoil sections (1) FX63-137, (2) NACA 4412, and (3) NACA 0012 are evaluated by using Fluent's software. The flow problems are treated as an internal flow problem with the assumption that the pressure ratio between the inlet and outlet of the air tunnel is 700 N/m2. The flow is treated as turbulent flow with k — ϵ as its turbulence modelling. In this respect Fluent uses a SIMPLE scheme for solving their governing equation of fluid motion. Their computational results on the way the object moves along the air tunnel, it had been the carrier which uses airfoil NACA 0012 has better performance than the carrier that uses airfoil FX 63-137 or airfoil NACA 4412. This research work gives contribution in understanding the movement of the carrier in the form wing-cube configuration inside an air tunnel.

The study presents observations on the interaction of double-blade propeller tip vortices with a smooth-wall turbulent boundary layer (TBL). The wall-bounded helicoidal vortices from the propeller modify the velocity profiles and turbulence statistics. The effects of two different tip clearances, $\epsilon = 0.1\delta _0$ and $0.5\delta _0$ , at a matched thrust, are explored with particle image velocimetry to understand the dynamics of tip-vortex formation within the logarithmic and wake regions of the boundary layer. The measurements are performed with $\lambda =U_{tip}/U_{\infty }$ in the range 5.3–5.9, and a blade passing frequency ( $\,f_{prop}$ ) of the same order of the boundary-layer time scale ( $\,f_{TBL}$ ). Observations indicate a reduction in the extent of the log region and an enhancement of the wake parameter $\varPi$ , mirroring the behaviour seen in TBLs under adverse pressure gradient conditions. Notably, the slipstream most contracted region exhibits a significant reduction in the skin friction coefficient $C_f$ and an amplification of the velocity fluctuation statistics across the entire boundary layer. At a clearance of $\epsilon = 0.1\delta _0$ , there is evidence of the formation of paired coherent wall-bounded structures. The presence of the wall decreases the amplitude of both periodic and stochastic fluctuations obtained with a phase-locked triple decomposition. An exception is observed behind the propeller for the stochastic fluctuations of the wall-normal component of the flow, which become amplified as the blades move away from the wall. This leads to the creation of a more intense phase-locked two-point spatial coherence than that observed in fluctuations aligned with the streamwise direction. Furthermore, results reveal that reduced tip clearances lead to higher viscous dissipation and more active energy exchange between the mean flow and organized motions.

A numerical investigation of low-Reynolds number flows with thermal effect around the MAV airfoils using various turbulence models, including an algebraic Baldwin-Lomax model, Spalart-Allmaras one equation, and two equation (k-ω and SST-k-ω) turbulence models, is presented. First, the thermal effect on the aerodynamic efficiency is studied for flow around a rectangular MAV wing, based on the NACA0012 airfoil section at low-aspect ratio (AR = 2) and an angle of attack equal to 0°. Second, details of the thermal effect are limited to the two-dimensional NACA0012 airfoil with chord length of 3.81 cm. This study shows that the improvement of aerodynamic efficiency (increase lift and reduce drag) is achieved by the generation of a temperature difference between extrados and intrados of the airfoil (by cooling the upper surface and heating the lower surface). The numerical results obtained with various turbulence models are in good agreement with experiment data, except the k-ω turbulence model. These results are performed with the CFD-FASTRAN code, using the fully implicit scheme for time integration and the upwind Roe flux difference splitting scheme for space discretization augmented by a high order Osher-Chakravarthy limiter.

The physics of a parallel blade―vortex interaction is studied numerically and the predicted pressure and acoustic results are compared with experimental measurements. A high-resolution solution of the compressible Euler equations is performed on structured overset meshes. Initially, a two-dimensional airfoil-vortex interaction is studied for both a case where the vortex misses the blade and a case of direct impact. The vortex is initiated in the flow as a perturbation to the freestream conditions and is free to evolve, thus allowing for the deformation of the vortex as it interacts with the blade to be studied. The simulation is seen to accurately reproduce the experimental results and the emission of the acoustic waves from the airfoil surface is observed in detail. Acoustic energy generated by the interaction is seen to primarily radiate from the leading-edge section of the airfoil with a weaker contribution coming from the trailing edge. The simulations are then extended to three-dimensional moving overset meshes where the vortex generation and convection is also resolved. The numerical methodology is seen to accurately preserve the vortex strength and accurately reproduce the experimentally measured blade surface pressures and acoustics. The computations presented here face similar challenges to that encountered in the simulation of realistic helicopter blade—vortex interaction, but the computational costs are such that the solutions can be well resolved, and comprehensively validated using moderate resources.