Showing papers on "Computational aeroacoustics published in 2001"

Wave computation in compressible flow using space-time conservation element and solution element method

Abstract: The recently developed space-time conservation element solution element method has several attractive features for aeroacoustic computations. The scheme is robust, possesses almost no dispersion error, and the implementation of nonreflecting boundary conditions is simple and effective. The scheme is tested for several problems ranging from linear acoustic waves to strongly nonlinear phenomena, with special emphasis on mixing-layer instability, and good numerical results are achieved.

Adaptive nonlinear artificial dissipation model for computational aeroacoustics

Abstract: An adaptive nonlinear artificial dissipation model is presented for performing aeroacoustic computations by high-order and high-resolution numerical schemes based on central finite differences. It consists of a selective background smoothing term and a well-established nonlinear shock-capturing term, which damps out spurious oscillations caused by the central differences in the presence of a shock wave and keeps the linear acoustic waves relatively unaffected. A conservative form of the selective background smoothing term is presented to calculate accurate propagation speed or location of the shock wave. The nonlinear shock-capturing term, which has been modeled by second-order derivative term, is combined with it to improve the resolution of discontinuity and enhance the numerical stability near the shock wave. An adaptive control constant for overall amplitude of the dissipation is automatically calculated according to given grid metrics and time-dependent flow conditions. It is shown that the improved artificial dissipation model reproduces the correct profile and speed of the shock wave, suppresses numerical oscillations near the discontinuity, and avoids unnecessary damping on the smooth linear acoustic waves. The feasibility and performance of the adaptive nonlinear artificial dissipation model for the computational aeroacoustics are investigated and validated by the applications to actual problems.

Multiresolution Wavelet Based Adaptive Numerical Dissipation Control for Shock-Turbulence Computations

Abstract: The recently developed essentially fourth-order or higher low dissipative shock-capturing scheme of Yee, Sandham and Djomehri (1999) aimed at minimizing nu- merical dissipations for high speed compressible viscous ows containing shocks, shears and turbulence. To detect non smooth behavior and control the amount of numerical dissipation to be added, Yee et al. employed an artifcial compression method (ACM) of Harten (1978) but utilize it in an entirely diferent context than Harten originally intended. The ACM sensor consists of two tuning parameters and is highly physical problem dependent. To minimize the tuning of parameters and physical problem depen- dence, new sensors with improved detection properties are proposed. The new sensors are derived from utilizing appropriate non-orthogonal wavelet basis functions and they can be used to completely switch off the extranumerical dissipation outside shock lay- ers. The non-dissipative spatial base scheme of arbitrarily high order of accuracy can be maintained without compromising its stability atall parts of the domain where the solution is smooth. Two types of redundant non-orthogonal wavelet basis functions are considered. One is the B-spline wavelet (Mallat & Zhong 1992 ) used by Gerritsen and Olsson (1996) in an adaptive mesh refinement method, to determine regions where refinement should be done. The other is the modification of the multiresolution method of Harten (1995) by converting it to a new, redundant, non-orthogonal wavelet. The wavelet sensor is then obtained by computing the estimated Lipschitz exponent ofacho- sen physical quantity (or vector) to be sensed on a chosen wavelet basis function. Both wavelet sensors can be viewed as dual purpose adaptive methods leading to dynamic numerical dissipation control and improved grid adaptation indicators. Consequently, they are useful not only for shock-turbulence computations but also for computational aeroacoustics and numerical combustion. In addition, these sensors are scheme inde- pendent and can be stand alone options for numerical algorithm other than the Yee et al. scheme.

Numerical prediction of the unsteady flow and radiated noise from a 3D lifting airfoil

Abstract: The numerical prediction of the aerodynamic noise radiated by an isolated airfoil is performed using a Computational AeroAcoustics (CAA) method. This hybrid method combines (i) a compressible three-dimensional Large Eddy Simulation (LES) of the nearfield unsteady flow and (ii) a Kirchhoff integration providing the noise radiated in the far field. This process is applied to a symmetrical NACA0012 airfoil with a constant section and a blunted trailing edge (TE), at a Mach number of 0.205 and an angle-of-attack of 5°. The Reynolds number based on the airfoil chord is 2.86 millions. The computational domain has a spanwise extent representing 3.3 % of the chord. The unsteady flow computed via LES exhibits a superimposed pressure fluctuation field which presents the qualitative and quantitative features of the TE noise generated by the acoustic scattering of (i) the turbulent boundary layers convected on both airfoil sides (broad-band noise) and (ii) the alternated vortex shedding generated by the TE bluntness (narrow band component). Due to the strong stretching of the LES computational grid, which acts as an acoustic low-pass frequency filter, this acoustic field cannot radiate farther than a half-chord from the body. Consequently, the LES must be relayed by an acoustic propagation method to correctly simulate the farfield noise, which is done using a Kirchhoff integral code. This method necessitates a careful parametric study of the position of the control integration surface enclosing the airfoil and its wake. Farfield noise predictions are compared with (published experimental data obtained in an anechoic facility with the same airfoil geometry (but much larger span). The overestimation of experimental data by predicted levels is discussed.

Effects of Forward Flight on Jet Mixing Noise from Fine-Scale Turbulence

Abstract: 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.

Finite-element method to study harmonic aeroacoustics problems

Abstract: The analysis of aeroacoustics propagation is required to solve many practical problems. As an alternative to Euler’s linearized equations, an equation was established by Galbrun in 1931. It assumes the flow verifies Euler’s equations and the perturbation is small and adiabatic. It is a linear second-order vectorial equation based on the displacement. Galbrun’s equation derives from a Lagrangian density and provides conservative expressions of the aeroacoustics intensity and energy density. A (CAA) method dealing with the numerical resolution of Galbrun’s equation using the finite-element method (FEM) is presented. The exact solution for the propagation of acoustic modes inside an axisymmetric straight-lined duct in the presence of a shear flow is known and compared with the FEM solution. Comparisons are found to be in good agreement and validate a first step in the development of a CAA method based on the FEM and Galbrun’s equation. The FEM is then applied to an axisymmetric duct including a varying cross...

Validation of a High-Order Prefactored Compact Code on Nonlinear Flows with Complex Geometries

Abstract: A finite-difference time domain solution of the airfoil gust problem is obtained using a high-accuracy nonlinear computational aeroacoustics code. For computational efficiency, the equations are cast in chain-rule curvilinear form, and a structured multiblock solver is used in parallel. In order to fully investigate the performance of this solver, a test matrix of eight problems are computed (two airfoil geometries, two gust frequencies, and two gust configurations). These results are compared to solutions obtained by the GUST3D frequency-domain solver both on the airfoil surface and in the far field. Grid density and domain size studies are included.

High Order Numerical Simulation of Sound Generated by the Kirchhoff Vortex

Abstract: An improved high order finite difference method for low Mach number computational aeroacoustics (CAA) is described. The improvements involve the conditioning of the Euler equations to minimize numerical cancellation errors, and the use of a stable non-dissipative sixth-order central spatial interior scheme and a third-order boundary scheme. Spurious high frequency oscillations are damped by a third-order characteristic-based filter. The objective of this paper is to apply these improvements in the simulation of sound generated by the Kirchhoff vortex.

Integral techniques for jet aeroacoustics calculations

Abstract: Reducing aircraft noise by a factor of four in the next twenty years is one of NASA's goals. Major reduction in acoustics emissions of aircraft jet engines is only possible with a reduction in jet noise. Hence, there is a need for improving the current state-of-the-art jet noise prediction methodology. We have developed in the past a Kirchhoff method code in order to evaluate the acoustic signals from an unsteady CFD code. In this paper, an integral acoustic code based on the porous Ffowcs Williams-Hawkings (FW-H) method is developed for the noise prediction of three-dimensional turbulent jets. The porous FW-H method can be more robust than the Kirchhoff method with regard to the choice of control surface, hence our efforts are focused in the development of the porous FW-H method. The resulting FWH code also includes refraction corrections to account for the zone of silence, as well as techniques to include acoustic sources outside the CFD domain. The code is validated for point sources. Once validation is complete, the code will be used to study the aeroacoustics of a three-dimensional turbulent jet at a Reynolds of number of 500 and possibly of another turbulent jet at a Reynolds number of 3600. The control surface needed for the FW-H method will be outside the jet flow, but inside the CFD boundaries in order to avoid boundary effects. The quantities on the control surface will be obtained using a Direct Numerical Simulation (DNS) code. The radiating sound will then be evaluated using the acoustics code developed herein.

Entropy Splitting for High Order Numerical Simulation of Vortex Sound at Low Mach Numbers

Abstract: Several recent developments in efficient, stable, highly parallelizable high order non-dissipative spatial schemes with characteristic based filters that exhibit low dissipation for long time linear and nonlinear wave propagations are utilized for computational aeroacoustics (CAA). For stability consideration, the Euler equations are split into a conservative and a symmetric non-conservative portion. Due to the large disparity of acoustic and stagnation quantities in low Mach number aeroacoustics, the split Euler equations are formulated in perturbation form to minimize numerical cancellation errors. Spurious oscillations are suppressed by a characteristic-based filter. The method has been applied to accurately simulate the sound emitted by an almost circular Kirchhoff vortex at low Mach numbers.

A Finite Volume Formulation for Compact Schemes on Arbitrary Meshes with Applications to LES

Abstract: In recent years there has been a great interest in compact (or Pade type) schemes [1], [2], [3]. On Cartesian uniform meshes these schemes offer a high order of accuracy with only a compact stencil. More importantly, the resolution of the shorter length scales of the solution of these schemes is better than for classical finite-difference methods [4], which makes these schemes especially attractive for applications such as DNS, LES and computational aeroacoustics [5], [6].

Entropy Splitting for High Order Numerical Simulation of Vortex Sound at Low Mach Numbers

Abstract: A method of minimizing numerical errors, and improving nonlinear stability and accuracy associated with low Mach number computational aeroacoustics (CAA) is proposed. The method consists of two levels. From the governing equation level, we condition the Euler equations in two steps. The first step is to split the inviscid flux derivatives into a conservative and a non-conservative portion that satisfies a so called generalized energy estimate. This involves the symmetrization of the Euler equations via a transformation of variables that are functions of the physical entropy. Owing to the large disparity of acoustic and stagnation quantities in low Mach number aeroacoustics, the second step is to reformulate the split Euler equations in perturbation form with the new unknowns as the small changes of the conservative variables with respect to their large stagnation values. From the numerical scheme level, a stable sixth-order central interior scheme with a third-order boundary schemes that satisfies the discrete analogue of the integration-by-parts procedure used in the continuous energy estimate (summation-by-parts property) is employed.

Automated Approach to Very High-Order Aeroacoustic Computations

Abstract: Computational aeroacoustics requires efficient, high-resolution simulation tools. For smooth problems, this is best accomplished with very high-order in space and time methods on small stencils. However, the complexity of highly accurate numerical methods can inhibit their practical application, especially in irregular geometries. This complexity is reduced by using a special form of Hermite divided-difference spatial interpolation on Cartesian grids, and a Cauchy-Kowalewski recursion procedure for time advancement. In addition, a stencil constraint tree reduces the complexity of interpolating grid points that am located near wall boundaries. These procedures are used to develop automatically and to implement very high-order methods (> 15) for solving the linearized Euler equations that can achieve less than one grid point per wavelength resolution away from boundaries by including spatial derivatives of the primitive variables at each grid point. The accuracy of stable surface treatments is currently limited to 11th order for grid aligned boundaries and to 2nd order for irregular boundaries.

Direct Calculations of Plume Dynamics of a Pulse Detonation Engine by the CE/SE Method

Abstract: This paper reports preliminary numerical results of plume dynamics for a pulse detonation engine (PDE). The space-time CE/SE method was used to solve the two-dimensional and axisymmetric Euler equations in conjunction with one chemical species equation in a time-accurate manner. Chemical reactions are simulated by a one-step, irreversible, finite rate model. The stiff source term in the species equation is treated by a point-wise implicit method based on a space-time volumetric integration of the source term over each CE. Numerical results show complex vortex/shock interactions in the vicinity of the thruster exit. Away from the PDE tube, spherical expansion of sound waves is evident. Although a much finer mesh will be needed to further resolve the relevant flow physics, the present results provide an encouraging first step toward the analysis of this challenging computational aeroacoustics problem.

Parametric Study of Weighted Essentially Nonoscillatory Schemes for Computational Aeroacoustics

Abstract: Based on the dissipation and dispersion relations, two groups of weighted essentially nonoscillatory (WENO) schemes are investigated for computational aeroacoustics. For high-order accuracy, all of the WENO schemes are required to be at least fourth-order accurate. Therefore, with a reasonable number of grid points, they can simulate the low-frequency waves with high-order accuracy. Also, for the higher frequency waves, two parameters in the WENO schemes are introduced to reduce their dissipation and dispersion errors. Both upwind and central schemes are considered and compared for order of accuracy and capability of wave capturing. A sequence of numerical simulations is carried out. These test problems are propagation of a sine-wave packet, propagation of a spherical wave, shock and sine wave interactions, and propagation of acoustic, vorticity, and density pulses in a freestream.

Optimized weighted Essentially Non-Oscillatory schemes for computational aeroacoustics

Abstract: ENO (Essentially Non-Oscillatory) and weighted ENO (WENO) schemes were designed for high-resolution of discontinuities such as shock waves. They are uniformly highorder accurate, and essentially oscillation free. Optimized schemes such as the DRP (Dispersion-Relation-Preserving) schemes are optimized for short waves (with respect to the grid spacing ∆x, e.g., waves that are 6-8∆x in wave length) in the wave number space. Although they may have formally lower order accuracy than non-optimized maximum-order schemes, they are capable of resolving short waves with higher accuracy. Therefore, they are better suited for broadband acoustic wave problems. In this paper, we seek to unite the advantages of ENO, WENO schemes and optimized schemes through the development of Optimized Weighted ENO (OWENO) schemes to tackle shock/broadband acoustic wave interactions and small scale flow turbulences relative to the grid spacing. A third-order OWENO and a seventh order WENO scheme are compared against each other for performance on the scalar model equation. It has been shown that the OWENO scheme indeed gives much better results in resolving short waves than the WENO scheme while yielding non-oscillatory solutions for discontinuities. The OWENO scheme is then extended to the linearized Euler equations to solve two computational aeroacoustics (CAA) benchmark problems for which analytical solutions are available. For both cases, excellent agreement with analytical solutions has been achieved.

Least-squares Legendre spectral element solutions to sound propagation problems.

Abstract: This paper presents a novel algorithm and numerical results of sound wave propagation. The method is based on a least-squares Legendre spectral element approach for spatial discretization and the Crank–Nicolson [Proc. Cambridge Philos. Soc. 43, 50–67 (1947)] and Adams–Bashforth [D. Gottlieb and S. A. Orszag, Numerical Analysis of Spectral Methods: Theory and Applications (CBMS-NSF Monograph, Siam 1977)] schemes for temporal discretization to solve the linearized acoustic field equations for sound propagation. Two types of NASA Computational Aeroacoustics (CAA) Workshop benchmark problems [ICASE/LaRC Workshop on Benchmark Problems in Computational Aeroacoustics, edited by J. C. Hardin, J. R. Ristorcelli, and C. K. W. Tam, NASA Conference Publication 3300, 1995a] are considered: a narrow Gaussian sound wave propagating in a one-dimensional space without flows, and the reflection of a two-dimensional acoustic pulse off a rigid wall in the presence of a uniform flow of Mach 0.5 in a semi-infinite space. The f...

High Order Difference Method for Low Mach Number Aeroacoustics

Abstract: A high order finite difference method with improved accuracy and stability properties for computational aeroacoustics (CAA) at low Mach numbers is proposed. The Euler equations are split into a conservative and a symmetric non- conservative portion to allow the derivation of a generalized energy estimate. Since the symmetrization is based on entropy variables, that splitting of the flux derivatives is referred to as entropy splitting. Its discretization by high order central differences was found to need less numerical dissipation than conventional conservative schemes. Owing to the large disparity of acoustic and stagnation quantities in low Mach number aeroacoustics, the split Euler equations are formulated in perturbation form. The unknowns are the small changes of the conservative variables with respect to their large stagnation values. All nonlinearities and the conservation form of the conservative portion of the split flux derivatives can be retained, while cancellation errors are avoided with its discretization opposed to the conventional conservative form. The finite difference method is third-order accurate at the boundary and the conventional central sixth-order accurate stencil in the interior. The difference operator satisfies the summation by parts property analogous to the integration by parts in the continuous energy estimate. Thus, strict stability of the difference method follows automatically. Spurious high frequency oscillations are suppressed by a characteristic-based filter similar to but without limiter. The time derivative is approximated by a 4-stage low-storage second-order explicit Runge-Kutta method. The method has been applied to simulate vortex sound at low Mach numbers. We consider the Kirchhoff vortex, which is an elliptical patch of constant vorticity rotating with constant angular frequency in irrotational flow. The acoustic pressure generated by the Kirchhoff vortex is governed by the 2D Helmholtz equation, which can be solved analytically using separation of variables.

Computational aeroacoustics of phonation: Effect of subglottal pressure, glottal oscillation frequency, and ventricular folds

Abstract: Numerical simulations of flow and acoustics in an idealized axisymmetric model of the human vocal tract have been conducted. The compressible Navier–Stokes equations were numerically integrated using highly accurate numerical treatments together with a forced glottal oscillation model and a moving grid. The effects of subglottal pressure and glottal oscillation frequency on velocity, vorticity, wall pressure, wall shear stress, and acoustic signal of the pulsating jet were investigated. An acoustic analogy was used to predict the far‐field sound. Excellent agreement between the predicted and directly simulated far‐field sound was obtained. The acoustic analogy was also used to decompose the acoustic source and identify monopole, dipole, and quadrupole contributions for analyses. The results show significant effects of subglottal pressure and glottal oscillation frequency on the jet vortical structure, wall forces, and acoustic radiation. The effect of including the ventricular (false) folds downstream of the oscillating glottal region was also investigated. Jet impingement on the ventricular folds introduces additional dipole sound sources. [Work supported by NIH DCO 3577‐02, RO1 grant from NICDC.]

Sound Emission of Rotor Induced Deformations of Generator Casings

Abstract: The casing of large electrical generators can be deformed slightly by the rotor''s magnetic field. The sound emission produced by these periodic deformations, which could possibly exceed guaranteed noise emission limits, is analyzed analytically and numerically. From the deformation of the casing, the normal velocity of the generator''s surface is computed. Taking into account the corresponding symmetry, an analytical solution for the acoustic pressure outside the generator is found in terms of the Hankel function of second order. The normal velocity of the generator surface provides the required boundary condition for the acoustic pressure and determines the magnitude of pressure oscillations. For the numerical simulation, the nonlinear 2D Euler equations are formulated in a perturbation form for low Mach number Computational Aeroacoustics (CAA). The spatial derivatives are discretized by the classical sixth-order central interior scheme and a third-order boundary scheme. Spurious high frequency oscillations are damped by a characteristic-based artificial compression method (ACM) filter. The time derivatives are approximated by the classical 4th-order Runge-Kutta method. The numerical results are in excellent agreement with the analytical solution.

Numerical Computation of Sound Propagation and Radiation in a Duct

Abstract: This paper presents an application of a computational aeroacoustics approach to the computation of sound propagation in a hard-wall duct and the associated radiation field including non-uniform mean flow effects. The linearized Euler equations are the base of the mathematical model. The dispersionrelation-preserving finite difference scheme is employed for space discretization whereas the low-dissipation and lowdispersion Runge-Kutta scheme is applied for time stepping. Appropriate non-reflecting far-field boundary conditions are prescribed at the inflow and outflow regions. Numerical results are given for sound wave propagaton in a transonic and supersonic nozzle. The sound wave propagation and radiation through a generic axisymmetric duct inlet is simulated with an input mean flow field calculated by a CFD code. Introduction High intensity acoustic waves are generated by compressors/fans inside an aero-engine and propagated out from the inlet and by-pass duct. The accurate prediction of acoustic mode propagaton in ducts and the radiation into the far-field has long been an important subject. Due to the potential to give more accurate and insightful predictions, such as accounting for mean flow, complex geometry and even the aerodynamic and acoustic coupling effects, numerical simulation using Euler or Navier-Stokes equations has gained favor by a lot of researchers in recent years. For example, Ozyoriik and Long[ll] utilized an Euler/Kirchhoff coupling method for the computation of sound radiation from engine inlets. This method has recently been extend to simulate fore and aft sound radiation from an engine geometry [1]. Both the acoustic and mean flow are based on the same governing equations. However, high order schemes designed for unsteady flow calculations are not so efficient for steady flow computations although some accelerating technique can be utilized. *Copyright©2001 by X.D. Li, N. Schonwald and F. Thiele Published by the American Institute of Aeronautics and Astronautics with Permission, t Associate professor, Department of Jet Propulsion, Member AIAA. * Graduate student, Hermann-Fo'ttinger-Institute for Fluid Mechanics. § Professor, Hermann-Fottinger-Institute for Fluid Mechanics. Stanescu, etc.[12] proposed a spectral method for computation of sound radiation from duct inlets. A bell-mouth duct is designed for investigating the mean flow effects. Both the mean flow field and acoustic field are solved by nonlinear Euler equations. This method was later applied to predict the sound radiation from the ONER A elbow and the JT15D turbofan inlet with no-mean-flow [2]. They suggested to start the acoustic calculation based on a mean flow field provided by another more efficient CFD code or from experiments. Although great advance has been achieved, a large number of validation test cases against analytical and experimental results have to be performed before a CFD or CAA code can be adopted for practical applications. This paper is to study the reliability and accuracy of calculating acoustic field using a linearized computational aeroacoustic (CAA) approach based on a known mean flow from CFD simulation or analytical solution. The 7 points dispersionrelation-preserving (DRP) scheme[14] is utilized for spatial discretization. The low-dissipation and low-dispersion Runge-Kutta scheme (LDDRK) [6] is applied for time integration. This basic CAA code has ever been validated by different benchmark problems chosen from the 1, 2 and 3 CAA workshops [9] [10]. In this paper, two benchmark problems from the third CAA workshop[19 are selected for validating the CAA procedure. One problem is on the propagation of sound waves passing through a transonic nozzle. Another problem is on shock/sound interaction in a supersonic nozzle. Finally a test case for a generic duct inlet is presented. The background mean flow field is calculated by a lower order Navier-Stokes solver. And the acoustic radiation is simulated based on the linearized Euler equations. Validation by Benchmark Problems In order to study the accuracy of the applied numerical schemes, two benchmark problems from the 3rd CAA workshop[16] were chosen for validation. The first problem is described as a ID sound wave propagation passing through a nearly chocked nozzle. The second problem is on a sound wave passing through a normal shock in a supersonic nozzle. (c)2001 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization. Governing Equations The governing equations of acoustics for both the problems are conservative quasi-ID linearized Euler equations:

New trends in turbulence = Turbulence : nouveaux aspects : Les Houches, session LXXIV, 31 July-1 September 2000

Abstract: The Century of Turbulence Theory: The Main Achievements and Unsolved Problems.- Measures of Anisotropy and the Universal Properties of Turbulence.- Large-Eddy Simulations of Turbulence.- Statistical Turbulence Modelling for the Computation of Physically Complex Flows.- Computational Aeroacoustics.- The Topology of Turbulence.- Burgulence.- Two-Dimensional Turbulence.- Analysing and Computing Turbulent Flows Using Wavelets.- Lagrangian Description of Turbulence.

Direct numerical simulation of 3-d jets using compact schemes and spatial filtering *

Abstract: Summary Direct numerical simulation (DNS) of threedimensional subsonic turbulent round jets is the main focus of this study. The full compressible threedimensional Navier-Stokes equations are solved using a parallel algorithm that employs a sixth-order compact scheme along with an eighth-order implicit spatial filter. Spatial filtering is used as the means of rejecting unwanted numerical oscillations that emanate from the boundaries of the computational domain. Tarn and Webb's non-reflecting boundary conditions are used to let acoustic waves freely exit the domain. Results for a turbulent jet of Reynolds number 1000 are presented. Excellent speedups up to 128 processors are obtained on an IBM SP machine. The current DNS code will be eventually used in a Computational Aeroacoustics (CAA) methodology for jet noise prediction.