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The Aircraft Noise Working Group (ANSWr) was established by DLR, ONERA, and NASA to compare simulation tools, establish guidelines for noise prediction, and to assess uncertainties associated with the simulation. To accomplish these goals, a benchmark problem was initiated by the group. The setup is documented and initial results for the reference aircraft are discussed. The reference aircraft is a conventional tube-and-wing configuration with the engines installed under the wings. The aircraft noise simulations are performed for departure and approach conditions, and the results obtained with three different system noise prediction tools are compared. At the aircraft level, the overall agreement is good between the three predictions. Peak noise levels agree within 3-4 dB. Fan and jet component predictions are generally very similar although there can be differences of up to 7 dB in some fan tone levels. The prediction of airframe components shows the most disagreement between the three methods with some differences of 6 dB for the major components and greater differences at high frequencies. There can also be differences in the frequency of the peak level and in the rank order of the airframe components. At the total airframe noise level, the differences are reduced to no more than 3-4 dB. While there is general agreement in shape characteristics of predicted ground noise isocontours, the differences between the three methods can result in more significant disagreement in the sizes of the isocontours., Le DLR, l’ONERA et la NASA ont créé le groupe de travail sur le bruit des avions (ANSWr) afin de comparer les outils de simulation, établir des directives pour la prévision du bruit et évaluer les incertitudes associées à la simulation. Pour atteindre ces objectifs, un problème de référence a été lancé par le groupe. La configuration est documentée et les premiers résultats pour l’avion de référence sont discutés. L'avion de référence est une configuration conventionnelle avec les moteurs installés sous les ailes. Les simulations de bruit des aéronefs sont effectuées pour les conditions de décollage et d’approche et les résultats obtenus avec les trois outils de prévision du bruit de chaque centre de recherche sont comparés. Au niveau des aéronefs, l’accord global est bon entre les trois prévisions. Les niveaux de bruit maximaux s'accordent avec un écart maximum de 3-4 dB. Les prévisions concernant le bruit des soufflantes et du jet sont généralement très similaires bien qu'il puisse exister des différences allant jusqu'à 7 dB entre certains niveaux de bruit tonal de soufflante. La prédiction du bruit aérodynamique de cellule montre le désaccord le plus important entre les trois méthodes, avec des différences de 6 dB pour les composants principaux et des différences plus importantes aux hautes fréquences. Il peut également y avoir des différences dans la fréquence du niveau maximum et dans l’ordre d'importance des différentes sources de bruit de cellule. Au niveau du bruit total de cellule, les différences ne sont pas supérieures à 3-4 dB. Bien qu’il existe un accord général sur les caractéristiques de forme des isocontours de bruit de sol prédits, les différences entre les trois méthodes peuvent donner lieu à un désaccord plus important concernant la taille des isocontours.

Liquid and slurry jets were injected through a circular orifice transverse to a M = 3.0 airflow. Mass samples of both jets were taken across the plume 30 injector diameters downstream. Pitot and static pressure surveys were taken across the liquid jet. These data allowed the calculation of distributions across the liquid jet plume of Mach number, air mass flow, liquid-to-air ratio, and momentum flux. A correlation for the liquid concentration in the downstream plane is also presented. In the plume, there is a core region of subsonic airflow carrying two-thirds of the mass collected in the plume. In the core, the liquid mass flow is nearly constant from side-to-side at a given height, and the average velocity of the liquid is only 30 to 60% of the local air velocity. A supersonic mixing region covering two-thirds of the area of the plume surrounds the core region. Comparison with the results from this direct sampling data indicate that correlations developed from photographic techniques are inadequate in determining the jet penetration and width of liquid and slurry jets. The slurry jet showed substantial phase separation. A 30% mass-loaded slurry of 1-5 µm silicon dioxide particles mixed with water was injected, and the local loading varied from a low of 13% at the bottom of the plume to 100% outside the liquid plume. The local loading increased as the jet boundary was approached from any direction. Master of Science

An innovative application of existing technology is proposed for attenuating the effects of transient phenomena, such as rotor-stator and rotor-strut interactions, linked to noise and fatigue failure in turbomachinery environments. A computational study was designed to assess the potential of Passive Porosity Technology as a mechanism for alleviating interaction effects and radiated noise by reducing the fluctuating forces acting on the vane surfaces. The study involved a typical high bypass fan stator airfoil immersed in a subsonic free field and exposed to the effects of a transversely moving wake. Time histories of the primitive aerodynamic variables obtained from Computational Fluid Dynamics (CFD) calculations were input into an acoustic prediction code to estimate noise levels at a radial distance of ten chords from the stator airfoil. This procedure was performed on the solid airfoil to obtain a baseline, and on approximately fifty porous configurations in order to isolate those that would yield maximum noise reductions without compromising the aerodynamic performance of the stator. It was found that, for a single stator immersed in a subsonic flow field, communication between regions of high pressure differential - made possible by the use of passive porosity - tends to induce a time-dependent oscillatory pattern of small inflow-outflow regions near the stator leading edge (LE), which is well established before wake effects come into play. The oscillatory pattern starts at the LE, and travels downstream on both suction and pressure sides of the airfoil. The amplitude of the oscillations seemed to be proportional to the extension of the porous patch on the pressure side. Regardless of this effect, which may not have occurred if the airfoil were placed within a stator cascade, communication between regions of high pressure differential is necessary to significantly alter the noise radiation pattern of the stator airfoil. Whether those changes result in noise abatement or enhancement depends primarily on the placement and extension of the porous patches. For most viable configurations, porosity reduced loading noise but increased thickness noise. Variations in nominal porosity were of secondary importance. In general, the best aerodynamic performers (i.e., those configurations that were able to reduce unsteady lift without severely altering the lift and/or drag characteristics of the solid airfoil) were also the best acoustic performers. As a result of using passive surface porosity, overall peak radiated noise was reduced by approximately 1.0 dB. This reduction increased to about 2.5 dB when the effects of loading noise alone were considered. Ph. D.

The study of normal impedance of perforated plate acoustic liners including the effect of bias flow was studied. Two impedance models were developed, by modeling the internal flows of perforate orifices as infinite tubes with the inclusion of end corrections to handle finite length effects. These models assumed incompressible and compressible flows, respectively, between the far field and the perforate orifice. The incompressible model was used to predict impedance results for perforated plates with percent open areas ranging from 5% to 15%. The predicted resistance results showed better agreement with experiments for the higher percent open area samples. The agreement also tended to deteriorate as bias flow was increased. For perforated plates with percent open areas ranging from 1% to 5%, the compressible model was used to predict impedance results. The model predictions were closer to the experimental resistance results for the 2% to 3% open area samples. The predictions tended to deteriorate as bias flow was increased. The reactance results were well predicted by the models for the higher percent open area, but deteriorated as the percent open area was lowered (5%) and bias flow was increased. A fit was done on the incompressible model to the experimental database. The fit was performed using an optimization routine that found the optimal set of multiplication coefficients to the non-dimensional groups that minimized the least squares slope error between predictions and experiments. The result of the fit indicated that terms not associated with bias flow required a greater degree of correction than the terms associated with the bias flow. This model improved agreement with experiments by nearly 15% for the low percent open area (5%) samples when compared to the unfitted model. The fitted model and the unfitted model performed equally well for the higher percent open area (10% and 15%). Ph. D.

A new unsteady rotor/fuselage interactional aerodynamics model has been developed. This model loosely couples a Generalized Dynamic Wake Theory (GDWT) to a Navier-Stokes solution procedure. This coupling is achieved using a newly developed unsteady pressure jump boundary condition in the Navier-Stokes model. The new unsteady pressure jump boundary condition models each rotor blade as a moving pressure jump which travels around the rotor azimuth =and is applied between two adjacent planes in a cylindrical, non-rotating grid. Comparisons are made between predictions using this new model and experiments for an isolated rotor and for a coupled rotor/fuselage configuration. Ph. D.

Reduction or cancellation of acoustic noise is achieved by providing an amplified, oppositely phased version of the noise by means of an acousto-fluidic amplifier. The amplified acoustic output noise is delivered through an impedance matching horn in destructively interfering relation with the original noise. Depending on the acoustic noise source and its spatial distribution, the acousto-fluidic amplifier may be a single stage amplifier or multiple stages connected in parallel and/or cascade, with output horns spatially distributed to have the maximum cancellation effect. Sensed noise, prior to fluidic amplification, may be processed in a manner to effect feedback or feedforward control of the amplified acoustic output signals.

An active noise control system using a compact sound source is effective to reduce aircraft engine duct noise. The fan noise from a turbofan engine is controlled using an adaptive filtered-x LMS algorithm. Single multi channel control systems are used to control the fan blade passage frequency (BPF) tone and the BPF tone and the first harmonic of the BPF tone for a plane wave excitation. A multi channel control system is used to control any spinning mode. The multi channel control system to control both fan tones and a high pressure compressor BPF tone simultaneously. In order to make active control of turbofan inlet noise a viable technology, a compact sound source is employed to generate the control field. This control field sound source consists of an array of identical thin, cylindrically curved panels with an inner radius of curvature corresponding to that of the engine inlet. These panels are flush mounted inside the inlet duct and sealed on all edges to prevent leakage around the panel and to minimize the aerodynamic losses created by the addition of the panels. Each panel is driven by one or more piezoelectric force transducers mounted on the surface of the panel. The response of the panel to excitation is maximized when it is driven at its resonance; therefore, the panel is designed such that its fundamental frequency is near the tone to be canceled, typically 2000-4000 Hz.

An active noise control system using a compact sound source is effective to reduce aircraft engine duct noise. The fan noise from a turbofan engine is controlled using an adaptive filtered-x LMS algorithm. Single multi channel control systems are used to control the fan blade passage frequency (BPF) tone and the BPF tone and the first harmonic of the BPF tone for a plane wave excitation. A multi channel control system is used to control any spinning mode. The multi channel control system to control both fan tones and a high pressure compressor BPF tone simultaneously. In order to make active control of turbofan inlet noise a viable technology, a compact sound source is employed to generate the control field. This control field sound source consists of an array of identical thin, cylindrically curved panels with an inner radius of curvature corresponding to that of the engine inlet. These panels are flush mounted inside the inlet duct and sealed on all edges to prevent leakage around the panel and to minimize the aerodynamic losses created by the addition of the panels. Each panel is driven by one or more piezoelectric force transducers mounted on the surface of the panel. The response of the panel to excitation is maximized when it is driven at its resonance; therefore, the panel is designed such that its fundamental frequency is near the tone to be canceled, typically 2000-4000 Hz.