Home HomeKrishnan K. Ahuja, Georgia Institute of Technology (U.S.A.)
Feedback phenomenon resulting in acoustic resonance is very common in a range of aeroacoustics problems of current interest. Screech in shock containing jets, cavity noise, edgetones, jet/collector interactions and howling of ejectors are but a few examples. In most of these problems involving shear layers, there is a match between the frequencies of sound and the most-preferred instability waves that are excited by the sound impinging at the edge where the shear layer begins. This paper will discuss the origins of this feedback phenomenon and the conditions under which it has the most impact. Selected examples of this phenomenon from the recent work conducted at Georgia Institute of Technology will be presented. Role of the boundary conditions and methods of controlling the phenomenon will also be discussed.
sv970537.pdf (Scanned)
TOPDonald B. Hanson, Pratt and Whitney (U.S.A.)
The problem of broadband noise generated by turbulence from a rotor impinging on a downstream stator is examined from a theoretical viewpoint. Equations are derived giving sound power spectra in terms of the 3 dimensional turbulence wavenumber spectrum. Particular attention is given to issues of turbulence inhomogeneity related to separation of individual blade wakes in the near field of the rotor. It is shown that this inhomogeneity can be handled rigorously with no additional complexity in the noise equations. A procedure for measurement of turbulence at a stator inlet with a 2 probe data system is studied. In the process, formulas are derived to compute the 3D turbulence spectrum via transforms of measured cross-spectra and estimates are given for the required probe spacing.
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M.K. Bull, University of Adelaide (Australia)
A.M. Blazewicz, University of Adelaide (Australia)
J.M. Pickles, University of Adelaide (Australia)
Acoustic radiation is generated by fluid flow over a tandem array of thick plates with bluff trailing edges, as a result of interaction between the leading edge of the downstream plate and vortices formed by flow separation from the upstream plate. The interaction gives rise to force fluctuations on the downstream plate, which constitute an acoustic dipole. The dominant component of the acoustic radiation is of discrete frequency, and the level is strongly dependent on the vortex flow regime which is established in the gap between the plates. The vortices may be stationary, trapped in the gap between the two plates, or may form a vortex street as a result of shedding from the upstream plate. Both flow regimes produce discrete frequency acoustic radiation, but with the onset of the vortex-street regime a dramatic increase in the level of acoustic radiation occurs. In broad terms, the flow regime - and hence the intensity of acoustic radiation produced - is determined by just two primary parameters: (i) the ratio S = s/c, of the lengths of any leading-edge separation bubble formed on the upstream plate to the chord-length c, of the plate, and (ii) the ratio G = g/t of the gap g between the plates to the plate thickness t. In general, the gap contains two trapped counter-rotating vortices or some form of vortex street, according as G > or < a critical value GC(typically - 3). For S <1, any leading-edge flow separation on the upstream plate is followed by reattachment on the same plate; flow separation from the trailing edge of the upstream plate then leads to a trapped-vortex or vortex-street regime in the gap (depending on the value of G). For S close to unity, the reattachment becomes intermittent and modifies both the trapped-vortex and vortex-street flow in the gap. For S > 1, leading-edge separation occurs without reattachment on the upstream plate, and the flow regime which is established depends also on the ratios s/(c1 + g) and s/(c1 + g + c2), where C2 is the chord-length of the downstream plate. Experimental data for a range of plate-array geometries will be presented to show that the various possible flow regimes are characterised by the coefficient of pressure on the upstream face of the downstream plate and by the Strouhal number of trapped-vortex oscillation or of vortex shedding from the upstream plate. The results clearly show that whenever the flow regime changes, these flow parameters and the level of the accompanying acoustic radiation all exhibit an abrupt change in magnitude.
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Kunisato Seto, Saga University (Japan)
Kiichi Tokuhisa, Saga University (Japan)
Muneharu Matsuoka, Saga University (Japan)
A Perforation in a wall over which air flows may be considered to have both noise generating and noise absorbing effects. The authors have experimentally investigated the effects of perforated wall with and without shroud or cavity on noise reduction. First, aerodynamic characteristics has been improved by arranging the perforation according to cross area change of a De Laval nozzle. Next a perforated tube has been combined with a shroud in order that no leakage occurs through the perforation and also acoustic transmission loss increases. Third the length of the shroud has been extended longer than the main tube to give it an effect of a cavity. The cavity containing tube has some cut-off frequency and augments turbulence mixing which is good for combustion and reduces noise in some case. For those three cases, the authors investigated the effects of the porosity, the size of the perforation, the intersection angle between the axes of the perforation and the main flow and the perforation pattern. The directional angle of the perforation has been found to have influence on the performance of the perforated tube.
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Stewart Glegg, Florida Atlantic University (U.S.A.)
William Devenport, Virginia Polytechnic and State University (U.S.A.)
This paper describes the use of proper orthogonal decomposition for problems in hydroacoustics where structural response and sound radiation is caused by unsteady flow. For a complete solution, the flow in the region of interest must be calculated using numerical solutions to the Navier Stokes equations and this paper will describe how proper orthogonal modes may be used to specify the unsteady flow at the upstream boundary of the computational domain. This will be shown to offer significant savings in computational effort for both linear and non-linear problems providing the correct modes can be dertermined.
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Ravi N Margasahayam, John F. Kennedy Space Center (U.S.A.)
Raoul E. Caimi, John F. Kennedy Space Center (U.S.A.)
This paper presents a brief overview of recently completed research in the area of rocket noise and resulting dynamic behavior of launch pad structures. To gain accurate insight into the vibratory behavior of these structures, dynamic tests were integrated into the design process. Aspects of the acoustic load characterization procedure and the test-analysis correlation of random vibration structural response in the low frequency range (1 to 50 hertz) are presented.
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Raoul E. Caimi, John F. Kennedy Space Center (U.S.A.)
Ravi N Margasahayam, John F. Kennedy Space Center (U.S.A.)
Jamal F. Nayfeh, University of Central Florida (U.S.A.)
This study consists of an investigation into the nonstationary transient response of the Verification Test Article (VETA) when subjected to random acoustic excitation. The goal is to assess excitation models that can be used in the design of structures and equipment when knowledge of the structure and the excitation is limited. The VETA is an instrumented cantilever beam that was exposed to acoustic loading during five Space Shuttle launches. The VETA analytical structural model response is estimated using the direct averaged power spectral density and the normalized pressure spectra methods. The estimated responses are compared to the measured response of the VETA. These comparisons are discussed with a focus on prediction conservatism and current design practice.
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Yoshiyuki Maruta, EBARA Research Co. Ltd. (Japan)
Sadao Mori, Metropolitan Expressway Public Corporation (Japan)
Hiroo Tanabe, EBARA Corporation (Japan)
It is necessary to estimate the exhaust-flow noise from ventilation tower when planning to increase the velocity of exhaust gas for preventing down-draft and down-wash around ventilation tower of vehicle tunnels. There are some aerodynamic noise source around the ventilating gas exit of vehicle tunnels, although many silencers are installed in ventilating duct for fully reducing the fan-noise. Small model experiments have been done by using the Quiet-Flow Acoustic Wind-Tunnel for studying the method of estimating aerodynamic noise radiated from ventilating gas exit of vehicle tunnels. In our experiments shape of exit port are rectangle and the length of exit-side were changed from 0.15m to 0.6m. The velocity of exhaust gas flow were changed from 10m/s to 30m/s.The aerodynamic noise power is proportional to the 6th power of gas-flow velocity and is proportional to the front side length of exit port. It is considered by these characteristics that the dominant noise source is aerodynamic noise generated by vortex shedding at the front side edge of exit port. Two-type equations with using side length of the exit for predicting exhaust flow noise have been induced by this experimental study. Basical methods to estimate the exhaust flow noise from ventilating gas exit of vehicle tunnels have been understood although there are still some subjects which need practical detail investigations.
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Rolf C. Pedersen, University of Western Australia (Australia)
Michael P. Norton, University of Western Australia (Australia)
Methods developed in the frequency domain are used for quantifying the hydrodynamic and acoustic components of the internal wall pressure fluctuations downstream of a flow disturbance in a flow duct system. This paper examines in detail, for frequencies below the (1,0) higher order acoustic mode, the acoustic and hydrodynamic fields for air flow downstream of a sharp edged orifice in a cylindrical flow duct. Particular attention is paid to the effect of the presence of a downstream standing wave field and the transmission of energy through the orifice.
sv970552.pdf (Scanned) TOPAlexander T. Fedorchenko, Moscow Phys. Tech. Institute (Russia)
New nonlinear theory of sound generated by unsteady subsonic flow has been set up. This two-medium theory is based on the nonlocal invariant procedure of separating out the acoustic components in high-unsteady flow and so it departs radically from all conventional approaches, including well-known Lighthills model. The theory also gives means for the simulation of sound generation by specified unsteady mass and heat sources as well as by body forces. Thus the proposed theory does change some universally recognized opinions in aeroacoustics. Nonlocal mathematical model has been created for the . simulation of unsteady subsonic flows, viscous or inviscid as required, in a bounded spatial domain under rather complex boundary conditions. Continuously distributed sources of mass and entropy as well as intensive heat conduction can be taken into account. The exclusion of all acoustic effects is an important feature of this model which represents an essential extension of the classical model of incompressible fluid flow. In turn, this model forms a necessary basis for further application of the above aeroacoustic theory. A number of topical problems has been studied with the use of both theoretical and computational methods: nonlinear processes of instability in free mixing layers and in internal viscous flows; interactions between coherent vortex structures and small-scale turbulence; sound generation and acoustic feedback; self-excited acoustic resonances in ducts of various shapes (with cavities, steps, sudden expansions, bends), etc. Effective means of control over bounded subsonic flows have been developed. The creating of high-stable nonuniform flows with minimum sound emission represents the main research goal that could promote a lot of practical applications.
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L.C. Chow, British Aerospace Airbus Ltd
R.J. Cummins, British Aerospace Airbus Ltd
The consideration of acoustically induced fatigue failures in aircraft has been a design consideration for many years. Acoustic fatigue failures can be caused by the dynamic response of aircraft structures to unsteady pressure loading from aerodynamic and engine acoustic sources. The life structures is often difficult to assess accurately and maybe greatly affected by steady, thermal, in-plane and out-of-plane panel loads. The failures can results in maintenance and inspection burdens associated with the operation of the aircraft. In the late 60s to mid 70s, industry and research establishments were involved in the development of dynamic response methods and generation of endurance data culminating in the publication of ESDU data sheets and AFFDL design guide. The information is still extensively used today. However, the airframe manufacturer is required to meet evermore stringent performance and mass targets, being achieved through developments in optimised and efficient design of structures, and the introduction of new materials, There is a corresponding need for development in acoustic fatigue design methods and data. A three year research project is currently in progress using advanced analytical procedure finite element analysis and complementary experimental studies in order to develop dynamic response prediction procedures which will result in guides for the design of box-type structures, such as flaps, composed of conventional aluminium alloys, CFRP and GLARE materials. This paper reports upon the experimental phase of the project and the results.
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S.R. Ahmed, DLR, Institute of Design Aerodynamics (Germany)
J.P. Yin, Nanjing University of Aeronautics and Astronautics (China)
W. Dobrzynski, DLR, Institute of Design Aerodynamics (Germany)
This numerical study investigates the influence of non-uniform rotation on the aerodynamics and acoustics of multi-blade propellers. Non-uniform rotational motion is inherent to piston engine driven propellers. The effect of rotational speed non-uniformity is the generation of excess harmonic noise due to unsteady aerodynamic blade-loading. In case of a mismatch between the periodicity of non-uniformity and the basic blade passing frequency, additional harmonics are generated due to the complex blade kinematics. For a periodicity coincidence the effects are masked due to overlapping of the frequencies. The level of such extra harmonics may be high enough to dominate the overall A-weighted noise level. Propeller noise radiation for non-uniform rotation is no longer omnidirectional along azimuth and exhibits also a different characteristic in the polar direction.
sv970420.pdf (Scanned) TOPJohan B.H.M. Schulten, National Aerospace Laboratory NLR (The Netherlands)
Of all outdoor noise sources, aircraft probably have the largest impact on communities. As a result, the accurate prediction of aircraft noise exposure is of great interest. Nevertheless, conventional procedures for quantifying aircraft noise draw heavily on empirical data in which source and propagation effects are more or less statistically lumped together. A physically more relevant modeling of aircraft noise propagation is the ray acoustics approximation. Whereas ray acoustics techniques are well developed for stationary sources, they are not often applied to aircraft noise because the aircraft motion in principle requires many time-consuming computations to obtain the time history of a single takeoff or landing event. The present paper describes the application of the method of ray-tracing to a source moving along a three-dimensional path in a realistic atmosphere. The method is illustrated by typical examples of the effects of a non-uniform wind and temperature profile such as the formation of acoustic shadow zones without any noise and, alternatively, zones with multiple reflections. It is shown that large reductions in computation time can be obtained if the flight path is close to level, which is factual for the majority of civil aircraft movements.
sv970393.pdf (Scanned) TOPThe paper examines the question: Does a mixing enhancement necessarily result in far-field noise reduction in supersonic jets? Several experiments were conducted to come up with an answer to the above question, The first experiment deals with examination of noise generated by a Mach 2 ideally expanded supersonic with counterflow. The counterflow technique depends on the creation of a counterflowing stream of air around the periphery of the primary jet column in the neighborhood of the nozzle exit. The creation of the secondary flow gives rise to a reduction in the jet exit plane pressure, which requires that the total pressure supplied to the jet be correspondingly reduced to maintain a nominally ideally expanded jet flow field. Experiments were conducted at a primary stream Mach number of 2 for jet stagnation temperatures between 286 K (cold jet) and 715 K (hot jet). The mixing characteristics of the jet were examined by conducting mean and fluctuating total pressure measurements in the axial and radial planes of the jet downstream of the nozzle-collar assembly. Fluctuating pressures measured along the geometrical axis of the jet indicate that the peak turbulence level lies closer to the jet exit when counterflow is applied. Complementary measurements made in the shear layer under similar flow conditions, indicate a significant increase in the overall turbulence level in the jet shear layers due to counterflow. These observations were essentially independent of the jet stagnation temperature. The corresponding mean total pressure measurements indicate that annular counterflow reduces the potential core length of the supersonic jets by a factor of two. Exhaustive studies conducted in our laboratory and summarized in Strykowski, Krothapalli & Jendoubi [1] attribule this mixing enhancement to increased shear layer growth rates by more than 60% compared to incompressible shear layers at similar velocity and density ratios. These aerodynamic measurements indicate that the counterflow significantly reduces the potential core, and therefore the supersonic region of the jet, and hence would appear to be an attractive control scheme for supersonic noise reduction. The most surprising result of the study was that the angle between the jet axis and the peak in the Overall Sound Pressure Level (OASPL) remained fixed when counterflow was supplied, despite the rather dramatic upstream shift in the peak turbulence level in the jet. This cast doubt on the notion that the primary source of the low frequency Mach wave radiation lies near the location of the peak rms pitot pressure. However, a closer examination of the measurements as outlined in King, Krothapalli & Strykowski [2] revealed that the reduction in the convection velocity of disturbances in the jet shear layer with counterflow essentially balanced the upstream shift in source location. Noise spectra obtained as a function of circular arc angle were used to explain why the OASPL of the counterflowing jet was higher at some angles relative to the free jet, but lower at others. In general, counterflow increases the noise levels at higher frequencies, suggesting that counterflow excites the smaller scale turbulence, a notion that is consistent with earlier findings that the rms pressure level in the jet shear layer was increased due to counterflow. At larger angles, the counterflow caused a reduction in the sound pressure level at all frequencies. Thus, the counterflow seems to interfere with the Mach wave radiation mechanism. In summary, these measurements suggest that enhanced mixing in the region close to the nozzle exit may not necessarily result in far-field noise reduction of a supersonic jet. This is primarily due to increased turbulence production that is closely associated with the enhanced mixing processes. The potential for using devices such as tabs and modifications to the nozzle exit geometry which may allow superior mixing through the introduction of streamwise vorticity without . significant thrust loss is explored. Although the mean and turbulent flow field of a jet at the nozzle exit with significant streamwise vorticity is very different than a corresponding axisymmetric jet , it does not appear to have a significant effect on the far field mixing noise of a supersonic jet. The flow and noise fields of a diamond-shaped jet are used to substantiate the above observation.
sv970346.pdf (Scanned) TOPKeiko Yoshida, Mitsubishi Electric Corporation (Japan)
The purpose of this study is to solve the vibration responses of spacecraft equipment structures, especially light weight rigid panels, such as a solar array paddle or an antenna, excited by an low frequency acoustical noise field. In this paper, first, the direct loading pressure to the structure in a random sound field, is derived and the response calculation method using FEM is proposed. Second, the derivation of the direct loading pressure to the structure is experimentally verified, using a honeycomb panel of which the edge was bolted to a stand, in the steady-state sound pressure field. Finally, experimental results of the acceleration responses of the panel in the random sound field are found to agree with the simulation results of this proposed calculation method. And the simulation results of the usual frequency response method contain a high degree of error compared with the results of the new method. This method is applied to a practical antenna panel, and the results of the simulation agree with the acoustic test.
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N.I Baranov, Russian Aviation Co. Lt (Russia)
S.N. Baranov, Russian Aviation Co. Lt (Russia)
Lev S. Kuravsky, Russian Aviation Co. Lt (Russia)
K.P Zhukov, Russian Aviation Co. Lt (Russia)
Computer experiments were undertaken to study dynamic behavior of a horizontal tail during the flight in turbulent atmosphere. This work was carried out to select acceptable variant of fixation and clearance between the stabilizer and elevator to avoid strikes.The method of correction by elastic and inertial connection insertion in Galerkin formulation was used to obtain a set of ordinary differential equations representing the structure motion. Under consideration is a sequence of mathematical models making each other more accurate. The following factors were successively taken into account: bending of the stabilizer and elevator, geometrical non-linearity of the structure response, effeot of the motion of a whole airplane, and torsion of the elevator. This approach made it possible to assess how important and correct each of these model modification for the estimations in question.To check the simulation correctness, computed results were compared with available experimental data obtained by ground vibration and flight testing. The estimations in question allowed aerospace designers to select optimal modifications of the horizontal tail.
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C.J. Gangadhara Gowda, Indian Institute of Technology (India)
Y.G. Srinivasa, Indian Institute of Technology (India)
Pramod S. Mehta, Indian Institute of Technology (India)
Control valves are the flow elements used in chemical process industries and steam power plants for regulating fluid flow in pipe lines. Control valve regulate the flow and convert pressure energy into kinetic energy. A portion of this energy is converted into acoustic energy which is responsible for generating noise. Studies on aerodynamic noise generated by control valves have attracted attention of several researchers for control of noise produced by them which is considerably high compared to the OSHA standards. The aerodynamic noise generated in control valve depends upon the pressure ratio between the upstream and downstream of the valve. For a higher pressure ratio in the system, the broad band shock noise is generated if the expansion is carried out in a single stage. This noise level is far greater than the noise level generated at sub-critical pressure ratio. The noise reduction in control valve, therefore can be achieved by allowing the fluid to expand in different stages such that the pressure ratio in each stage is kept below critical pressure ratio. The internal peak frequency of the noise generated by the control valve depends upon the jet diameter. Multi-hole expansion arrangement shifts the internal peak frequency depending upon jet diameter. This shift may be achieved such that the peak frequency occurs beyond the audible range hence reduce the noise level. The ability of the multi-hole and multi-stage arrangements to reduce noise can be combined for use in control valve trim design. From these considerations, the present study is an attempt to predict the aerodynamic noise generation from multi-hole multi-stage arrangement. In the present work the noise generated by a (i) multi-stage and (ii) multi-hole multi-stage expansion in a control valve are predicted for various pressure ratios. The results are compared with the predicted value of noise generated by a conventional control valve. There is considerable reduction in noise level when a multi-hole multi-stage configuration is used.
sv970280.pdf (Scanned) TOPChristopher K. W. Tam, Florida State University (U.S.A.)
Reducing high-speed jet noise is currently a high priority research and development item of the aircraft industry, The objective of this paper is to examine the effectiveness of jet noise reduction by nozzle geometry modification, The use of nozzles with complicated geometry would, inevitably, lead to thrust loss, Our consideration is, therefore, confined to practical geometries for which the thrust loss is reasonably small, In order to focus attention on nozzle geometry alone, we will only consider jets with a single stream. Multi-stream jets, invariably, would introduce thermodynamic and other flow parameters as variables for noise suppression, We have carried out a detailed analysis of the noise data from 6 nozzles operating at supersonic conditions, The nozzle geometry includes a CD configuration, a conical nozzle, a CD plug nozzle, a conical plug nozzle and two 20 chute annular plug suppressor nozzles, The results indicate that if the jet fluids mix quickly, that is, they combine into a single jet at a short distance downstream of the nozzle exit, the noise intensity and spectral characteristics are nearly the same as those of a simple round jet. There is a strong implication that external mixing introduced by nozzle geometry modification is not an effectiveness way to reduce jet noise.
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Sergey Ju. Krasheninnikov, Central Institute of Aviation Motors (Russia)
Alexei Mironov, Central Institute of Aviation Motors (Russia)
Eugeni V. Pavlyukov, Central Aerohydrodynamic Institute TsAGI (Russia)
Vladimir Jitenev, Central Aerohydrodynamic Institute TsAGI (Russia)
Andrey V. Shenkin, Central Aerohydrodynamic Institute TsAGI (Russia)
In earlier work of CIAM/TsAGI/SNECMA the acoustic and aerodynamic characteristics of 2-D mixer-ejector models, which could be prototype of real variable nozzle for future supersonic aircraft, were experimentally investigated. It was shown that 2-D mixer-ejector can give about 10 EPNdB noise reduction in take-off and flight over check points. Its efficiency is not worse than for conventional noise suppressors - about 34.5PNdB per 1% thrust loss at static conditions. In the present work the possibility of the increasing of the efficiency of the previously designed mixer-ejector by means of variation of its elements was studied. The results of the investigation of the influence of auxiliary slots and acoustic lining at ejector walls on mixer-ejector acoustic and thrust performances are presented. The flow structure and its influence on mixer-ejector acoustic characteristics was also studied.
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Dieter Lohmann, DLR Inst. of Design Aerodynamics (Germany)
H. Capdevila, Siemens Electric Ltd (Germany)
U. Stark, Technical University of Braunschweig (Germany)
M. Kuntz, DLR Inst. of Design Aerodynamics (Germany)
The need of aeroacoustically highefficient fans requires more detailed investigations of three dimensional effects and separated flow analysis. Comprehensive studies and code validation for fans have shown, that further development - especially for the influence of sweep - is necessary. Therefore, two different Navier-Stokes Codes, a lifting surface code and a classical method for 3-D cascade flow were used to compute a ducted swept radiator fan. The results of the method are checked against each other and compared with experiments to firstly validate the codes in aerodynamis and secondly deliver data for acoustic calculations. Here the aerodynamic pressure and boundary displacement thickness are taken for the acoustic field computations at the fan face using the acoustic analogy method. The far-field is computed by the use of Rayleigh's formula. Different flow rates were modeled for a 7-bladed fan geometry. The computational results of static pressure and efficiency are compared with experimental data. Indicated flow separation of both, a forward and an aft swept fan have been analysed. The effect of turbulence modeling and different grid size has been investigated. The linear methods used for the aeroacoustic fan design are validated for local data. In order to check the prediction capability of the DLR acoustic methods, acoustic spectra of periodic and stochastic noise parts have been calculated at different angles around the fan face and compared with measured data.
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Tetsu Yamagishi, University of Tokyo (Japan)
Shojiro Kaji, University of Tokyo (Japan)
The screech tone generated by a two-dimensional supersonic jet is studied numerically by solving the Euler equations with the ENO (essentially non-oscillatory) scheme. The details of shock cell structure, large-scale vortices, and sound generation mechanisms are elucidated. Numerical results for streak lines show clearly that shear layers begin to roll up at the third shock cell and grow into large-scale vortices destroying the shock cell structure further downstream. Beyond the third shock cell large-scale vortices cease to grow and convect at a nearly constant speed. In the growing process of large-scale vortices, it is observed that the jet plume and the vortex core compress the ambient air confined between them increasing its total temperature, whereas both of them losing their own energy. This higher energy fluid is pressed out of the vortex structure forming a sound source. Acoustic intensity analysis indicates that several sound sources along the jet plume bring acoustic interactions which make it possible that an acoustic energy feedback occurs only at the first shock cell.
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G. Guidati, University of Stuttgart (Germany)
Siegfried Wagner, University of Stuttgart (Germany)
The sound generation due to convected turbulence (vorticity) interacting with the leading edge of an arbitrary airfoil was treated by GUIDATI, ET AL [1]. In this approach the spectral decomposition of a point vortex being passively convected by the mean flow results in vorticity waves which are taken as basic source patterns. The acoustic analogy by HOWE [2] is used and the diffraction problem is solved by employing boundary-element methods. Comparison with recent experiments shows that the difference due to airfoil shape is predicted within an accuracy of 1-2 dB. In the present paper this approach will be extended. So far, only one vorticity wave on both sides is taken. Since a turbulent flow-field is completely filled with vorticity, a larger number will be employed. This allows to prescribe skewed gusts, ie turbulent length scales in flow and cross-flow direction. A second extension concerns the interaction of inflow turbulence with the trailing edge. In the AMIET model [3] a steady KUTTA condition is applied at the trailing edge and the development of a turbulent boundary layer is neglected. In reality, the question of a correct KUTTA condition is unsolved. Furthermore, the boundary layer will radiate additional sound and modify the inviscid flow field. Simulations with and without steady KUTTA condition will be presented. The deflection of streamlines due to a boundary layer will be considered. This will allow to estimate which part of the airfoil - leading or trailing edge - radiates more sound in case of turbulent inflow. Comparison with experiments will be shown.
sv970090.pdf (Scanned) TOPFritz-Reinhard Grosche, German Aerospace Research Establishment (DLR) (Germany)
Flow visualization techniques were used to study the influence of sound waves on flow separation from a wing of low aspect ratio 3:1. Two different technique of acoustic excitation have been applied: 1. Spherical sound waves from a loudspeaker within the acoustic far field (Global excitation) 2. Sound waves focused by an elliptical mirror on a small region of the wing (Localised excitation). The experimental results demonstrate that acoustic excitation in a suitable frequency range can reduce the flow separation, which occurs at high angles of incidence all along the leading edge, to a much smaller turbulent separation region that mainly affects the central part of the wing. The tests with acoustic waves focused on a small part of the wing give insight into the dependence of the spanwise structure of the separation on, the position of the excitation region. We made video-recordings of the flow visualization studies which illustrate the characteristic flow features obtained with global and localised acoustic control of the leading edge separation. Also presented are experiments on the control of separation by internal acoustic forcing through a slot in the surface of the model.
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F. de la Iglesia, Universidad Polictecnica de Madrid (Spain)
P. Garcia-Fogeda, Universidad Polictecnica de Madrid (Spain)
A method to estimate the response of an aerospace structure excited by the acoustical loads produced during a rocket launch is presented. These elevated pressure loads can be critical in the design of large lightweight structures such as solar arrays and communication reflectors where high acceleration levels can be achieved. The resulting acoustic field can be considered as a diffuse field composed of a large number of uncorrelated incident plane-waves traveling in different directions that impinge over the structure surface. A Boundary Element Method has been used to compute the pressure jump produced by an incoming plane-wave in an unbaffled rectangular plate and the fluid-structure coupling loads generated by the plate own vibrations. This method is based on Kirchhoffs integral formulation of the Helmholtz equation for the pressure field taking into account the Sommerfeld radiation condition. The generalized forces matrix due to the fluid loading is then determined utilizing the vacuum modes of the plate as base functions of the structural displacement in the present problem. These modes are obtained by means of a Finite Element Model. An iteration procedure has been developed to calculate the natural frequencies of the coupled fluid-plate system. Comparisons of the present method with various experimental data and other theories show the efficiency and accuracy of this method for any support condition of the plate and the validity of the present procedure for the values of the frequency of excitation that appears in an acoustical test performed in a large reverberant chamber.
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