Nikolai N. Pastouchenko

Fine-Scale Turbulence Noise from Hot Jets

Experimental measurements indicate that the noise radiated from a jet depends not just on the jet-exit velocity alone, but is significantly affected by the jet temperature. Now, there is evidence to support the proposition that jet mixing noise consists of two principal components. These are the noise from the large turbulence structures of the jet flow and the fine-scale turbulence. The prediction of fine-scale turbulence noise from hot jets is considered. Earlier Tam and Auriault developed a semi-empirical theory capable of predicting the fine-scale turbulence noise from cold to moderate temperature jets. In this work, their semi-empirical theory is extended to high-temperature jets, up to a temperature ratio above that of present day commercial engines. The density gradient present in hot jets promotes the growth of Kelvin-Helmholtz instability in the jet mixing layer. This causes a higher level of turbulent mixing and stronger turbulence fluctuations. In addition, recent experiments reveal that the two-point space-time correlation function of turbulent mixing for hot jets is substantially different from that for cold jets. The eddy decay time is shorter, and the eddy size is slightly reduced. These changes have an appreciable impact on the noise radiated. In the present extended fine-scale turbulence theory, both effects are taken into account

Noise from fine-scale turbulence of nonaxisymmetric jets

The noise from the fine-scale turbulence of high-speed nonaxisymmetric jets is considered. A prediction method is developed by extending the work of Tam and Auriault. A set of improved numerical boundary conditions for use in nonaxisymmetricjet mean flow and turbulence calculation is developed. These new boundary conditions allow the computation to be carried out in a smaller computation domain. It is known that nonaxisymmetric mean flow has a significant impact on the radiated noise spectrum and directivity through refraction. In the Tam and Auriault theory, this effect is accounted for by means of the adjoint Greens function. Here the adjoint Greens function method is extended to nonaxisymmetric mean flows. The adjoint Greens function is first recast into the solution of a sound scattering problem. The sound scattering problem is then solved computationally by computational aeroacoustics methods. Extensive comparisons between calculated and experimentally measured jet noise spectra are presented. They include both rectangular and elliptic jets at supersonic and subsonic Mach numbers

Effects of forward flight on jet Mixing Noise from fine-scale turbulence

It is known experimentally that a jet in forward flight radiates less noise than the samejetin a static environment. At a forward flight Mach number of 0.2, the noise reduction, depending on the jet operating conditions, could be as large as 4-5 dB in the sideline directions. In the past, a way to predict flight effects was to use the method of relative velocity exponent. Another method was to extrapolate measured static jet noise to the flight condition by means of scaling formulas. Both methods are semi-empirical. The fine scale turbulence jet mixing noise theory of Tam and Auriault (Tam, C. K. W., and Auriault, L., Jet Mixing Noise from Fine Scale Turbulence, AIAA Journal, Vol. 37, No. 2, 1999, pp. 145-153) is extended for application to jets in simulated forward flight. It will be shown that the calculated noise spectra at different simulated forward flight Mach numbers for both supersonic and subsonic jets compare well with experiments. The effects of forward flight on the sources of fine-scale turbulent jet mixing noise is also investigated. It is found that in the presence of forward flight the dominant noise sources move downstream. The turbulence intensity and the size of turbulent eddies responsible for noise emission are reduced.

Combustion Noise of Auxiliary Power Units

Noise from auxiliary power units (APU) is an important contributor to the overall level of ramp noise. Currently, ramp noise is regulated by international governing bodies as well as by individual airport. A significant component of APU noise is combustion noise. In this study, the unique spectral shape of APU combustion noise is identified. It is found that the spectral shape is the same regardless of engine size, power setting and directivity. Also, it is practically the same as that of open flame combustion noise. The frequency at the peak of the combustion noise spectrum is found to lie in the narrow range between 250 to 350 Hz. The peak sound pressure level of a given APU varies as the square of the fuel consumption rate. In the literature, suggestions have been made concerning a second combustion noise mechanism arising from the passage of hot entropy spots through the exhaust nozzle or constriction. In this investigation, no evidence has been found to indicate the existence of a second APU combustion noise component.

Fine-Scale Turbulence Noise from Dual-Stream Jets

Nowadays, commercial aircrafts, invariably, use high-bypass-ratio dual-stream jets for propulsion. As yet, there is still an urgent need for an accurate physics-based noise prediction theory for jets of this configuration. Thus, an investigation is made to determine whether the Tam and Auriault theory, originally developed for predicting the fine-scale turbulence noise of single-stream jets, is capable of predicting accurately the fine-scale turbulence noise of dual-stream jets from separate flow nozzles operating at various bypass ratios. The configuration of a separate flow nozzle is fairly complex. Hence, the jet flow and turbulence in the nozzle region and in the region immediately downstream are also fairly complex. However, these are also the most important noise source regions of the jet. To enable an accurate computation of the mean flow and turbulence level in these regions, a computational aeroacoustics marching algorithm for calculating the parabolized Reynolds averaged Navier-Stokes equations supplemented by the k-e turbulence model is provided

Experimental Validation of Numerical Simulations for an Acoustic Liner in Grazing Flow

A coordinated experimental and numerical simulation effort is carried out to improve our understanding of the physics of acoustic liners in a grazing flow as well our computational aeroacoustics (CAA) method prediction capability. A numerical simulation code based on advanced CAA methods is developed. In a parallel effort, experiments are performed using the Grazing Flow Impedance Tube at the NASA Langley Research Center. In the experiment, a liner is installed in the upper wall of a rectangular flow duct with a 2 inch by 2.5 inch cross section. Spatial distribution of sound pressure levels and relative phases are measured on the wall opposite the liner in the presence of a Mach 0.3 grazing flow. The computer code is validated by comparing computed results with experimental measurements. Good agreements are found. The numerical simulation code is then used to investigate the physical properties of the acoustic liner. It is shown that an acoustic liner can produce self-noise in the presence of a grazing flow and that a feedback acoustic resonance mechanism is responsible for the generation of this liner self-noise. In addition, the same mechanism also creates additional liner drag. An estimate, based on numerical simulation data, indicates that for a resonant liner with a 10% open area ratio, the drag increase would be about 4% of the turbulent boundary layer drag over a flat wall.

Computation of Shock Cell Structure of Dual-Stream Jets for Noise Prediction

Broadband shock cell noise is an important component of aircraft interior noise during cruise. At cruise, the secondary jet of most modern day jet engines is supersonic. As a result, a shock cell structure develops in the jet plume. The interaction of large turbulence structures of the jet flow and the periodic components of the shock cells results in the emission of highly directional broadband shock cell noise. The primary objective of this investigation is to develop a computational method to calculate the shock cell structures ofdual-stream jets issued from separate flow nozzles. Computation of the Fourier modes of such shock cell structures is also considered. Based on the dominant wave numbers computed by the method developed, the frequencies at the peaks of broadband shock cell noise spectra at various radiation directions are calculated. Good agreements are found with experimental measurements over a wide range of primary and secondary jet Mach numbers. The good agreements provide a validation of the accuracy of the computation method.

Continuation of Near-Acoustic Fields of Jets to the Far Field: Part II Experimental Validation and Noise Source Characteristics

This is the second of a two-part paper. This part reports the results of a systematic experimental study of the near acoustic field of a high-speed jet. The aim is to shed light on the noise source characteristics of the jet. A microphone array consisting of twenty-one semicircular rings of microphones is built for the experiment. The microphone array extends from one diameter from the nozzle exit to a distance of thirty-one diameters downstream. The microphones lie on a conical surface of ten degrees half-apex angle surrounding the jet. Two-point space-time pressure correlations are measured. There are two principal objectives in this study. The first objective is to validate the near field to far field continuation method, developed in the Part I paper, with experimental data. For this purpose, two series of validation tests are performed. First is a comparison of far field noise spectra between those continued from near field measurements on the conical surface and those measured directly in the far field. Another test series consists of comparisons of far field directivities at selected Strouhal numbers, again, between those continued from the near field measurements to the far field and those measured directly in the far field. Favorable agreements are found. These tests not only validate the continuation method developed in the Part I paper but also support the recognition that the two-point space-time pressure correlation function on the conical surface is the equivalent source of jet noise. This realization offers a way to investigate the characteristics of jet noise sources by analyzing those of the equivalent noise sources. The second objective of this experimental study is to investigate the noise source location, strength, size, convection speed, degree of randomness and other characteristics along the length of the jet. The noise source can also be decomposed into azimuthal modes. It will be shown that the axisymmetric noise source mode is the most dominant. Contributions to the radiated noise from modes higher than the second mode are negligible. Other modal characteristics such as propagation speed, location of maximum intensity and others will be reported. Finally, strong evidence will be presented to show the wave-like nature of the source of high-speed jet noise.

Installation effects on the flow and noise of wing mounted jets

T IS known experimentally, since the late 1970s, that a jet installed under a wing of an aircraft radiates more noise than the same jet in a standalone condition. The excess noise is the propulsion-airframe integration noise or commonly referred to as installation noise. When a jet is placed near a wing, there is an increase in noise in the flyover directions because of the reflection of sound by the wing. Here, installation noise includes not merely the noise increase due to the reflection of sound by the wing. The major part of this noise is generated aerodynamicly by the nonlinear interaction between the flow around the wing flap and the jet. In this work,ourprimaryinterestistomodelandtopredictinstallationnoise of aerodynamic origin. Installation noise increases not only the total aircraftnoiseinthe flyoverplanebutalsointhesidelinedirections.It is especially important during landings and takeoffs when the flaps are down. Duringthe1980s,anumberofexperimentswerecarriedouttrying to quantify the characteristics and intensity of installation noise [1– 4]. Most of these experiments involved the measurements of the jet alone noise and the noise when the jet was placed near a model of an aircraft wing inside an anechoic chamber. The experimental measurements by Wang [2] were the most systematic. In his experiment,ascaledmodelofthewingofaDC-10aircraftwasused. Large noise increase was observed in the flyover plane in the lowfrequencypartofthespectrum.Theincreaseinhigh-frequencynoise was less. In directions at small exhaust angles, the installation noise intensity was quite low. In the sideline, the radiated noise characteristics, on the other hand, were quite different. Overall, the measured data indicated that installation noise had a unique spectral shape and a directional pattern of its own. Recently, there is a renewed interest in propulsion-airframe integration noise. Mead and Strange [5] investigated the under-thewing installation effects on jet noise with special emphasis on the sidelinedirections.Theirinterestinthesidelinewasmotivatedbythe experience that it was generally more difficult to meet legislative limit on sideline noise level requirements. They reported the measurement of high installation noise level in the low-frequency range. One drawback of the Mead and Strange experiment [5], as well as most of the previous works, is that the experiments were carried out in static conditions. Upon realizing that the effect of forward flight is extremely important in the interaction between the flow around the wing flapandthejet,aseriesofnewexperimentsoninstallationnoise was conducted by engineers of The Boeing Company (Shivashankara and Blackner [6], Blackner and Bhat [7], and Bhat and Blackner [8]). They employed an open wind tunnel at M � 0:28 tosimulatetheforwardmotionoftheaircraft.Byusingellipticmirror microphones and the newly developed phase-array microphones, they were able to obtain noise source location maps as well as farfield noise data.

Numerical simulation of grazing incidence of sound waves on an acoustic liner

Acoustic liner is extremely effective for suppressing fan noise of jet engines. A resonant acoustic liner consists of a face sheet with cavity backing. Numerous small holes are drilled on the face sheet. When a sound wave is incident on a liner, pressure on the liner surface alternates from high to low. At high pressure, fluid is forced into the liner cavities through the holes. At low pressure, the process is reversed. The oscillatory motion of the fluid masses at the hole-openings is crucial to the damping of the sound waves. However, the hole diameter is typically one millimeter or less. Because the holes are small, experimental measurements of the fluid motion around the hole-openings are difficult to perform. Hence, this task is best carried out by numerical simulation; as small holes are not detrimental to numerical computation. The objective of this investigation is to seek an understanding of the flow physics responsible for acoustic damping by a liner. In all previous investigations, only one res...