Bart A. Singer

COMPUTATIONAL AEROACOUSTIC ANALYSIS OF SLAT TRAILING-EDGE FLOW

An acoustic analysis based on the Ffowcs Williams and Hawkings equation was performed for a high-lift system. As input, the acoustic analysis used unsteady flow data obtained from a highly resolved, time-dependent, Reynolds-averaged Navier-Stokes caclulation. The analysis strongly suggests that vortex shedding from the trailing edge of the slat results in a high-amplitude, high-frequency acoustic signal, similar to that which was observed in a corresponding experimental study of the high-lift system.

A predictor-corrector technique for visualizing unsteady flow

Presents a method for visualizing unsteady flow by displaying its vortices. The vortices are identified by using a vorticity-predictor pressure-corrector scheme that follows vortex cores. The cross-sections of a vortex at each point along the core can be represented by a Fourier series. A vortex can be faithfully reconstructed from the series as a simple quadrilateral mesh, or its reconstruction can be enhanced to indicate helical motion. The mesh can reduce the representation of the flow features by a factor of 1000 or more compared with the volumetric dataset. With this amount of reduction, it is possible to implement an interactive system on a graphics workstation to permit a viewer to examine, in 3D, the evolution of the vortical structures in a complex, unsteady flow. >

Time-Accurate Simulations and Acoustic Analysis of Slat Free-Shear Layer

Unsteady computational simulations of a multi-element, high-lift configuration are performed. Emphasis is placed on accurate spatiotemporal resolution of the free shear layer in the slat-cove region. The excessive dissipative effects of the turbulence model, so prevalent in previous simulations, are circumvented by switching off the turbulence-production term in the slat cove region. The justifications and physical arguments for taking such a step are explained in detail. The removal of this excess damping allows the shear layer to amplify large-scale structures, to achieve a proper non-linear saturation state, and to permit vortex merging. The large-scale disturbances are self-excited, and unlike our prior fully turbulent simulations, no external forcing of the shear layer is required. To obtain the farfield acoustics, the Ffowcs Williams and Hawkings equation is evaluated numerically using the simulated time-accurate flow data. The present comparison between the computed and measured farfield acoustic spectra shows much better agreement for the amplitude and frequency content than past calculations. The effect of the angle-of-attack on the slats flow features radiated acoustic field are also simulated presented.

Vortex tubes in turbulent flows: identification, representation, reconstruction

A new algorithm for identifying vortices in complex flows is presented. The scheme uses both the vorticity and pressure fields. A skeleton line along the center of a vortex is produced by a two-step predictor-corrector scheme. The technique uses the vector field to move in the direction of the skeleton line and the scalar field to correct the location in the plane perpendicular to the skeleton line. With an economical description of the vortex tubes cross-section, the skeleton compresses the representation of the flow by a factor of 4000 or more. We show how the reconstructed geometry of vortex tubes can be enhanced to help visualize helical motion.<<ETX>>

Reynolds-averaged Navier-Stokes computations of a flap-side-edge flowfield

An extensive computational investigation of a generic high-lift configuration comprising a wing and a half-span flap reveals details of the mean flow field for flap deflections of 29 and 39 degrees. The computational effort involves solutions of the thin layer form of the Reynolds Averaged Navier-Stokes(RANS) equations. For both flap deflections, the steady results show the presence of a dualvortex system; a strong vortex forming on the lower portion of the flap side edge and a weaker one forming near the edge on the flap top surface. Downstream, the vortex on the flap side edge grows and eventually merges with the vortex on the flap top surface. Comparison of on- and off-surface flow quantities with the experimental measurements of Radeztsky, Singer and Khorrami (AIAA Paper 98-0700) show remarkable agreement. For the 39 degree flap deflection, the calculation also reveals the occurrence of a vortex breakdown, which is corroborated by 5-hole probe velocity measurements performed in the Quiet Flow Facility at NASA Langley. The presence of the vortex breakdown significantly alters the flow field near the side edge.

Detailed measurements of a flap side-edge flow field

Detailed flow measurements were performed with a five-hole probe in the flap side-edge region of a wing with a half-span flap. The experiments were conducted in the NASA-Langley Quiet Flow Facility as a part of an extensive experimental and computational investigation of airframe-related noise sources. Basic flow properties were verified using static pressure taps, pressure-sensitive paint, and oil-flow visualization. Detailed five-hole probe surveys were conducted in a set of planes at ten streamwise stations for flap deflections of 29 and 39 degrees. These measurements show the development of a two-vortex system at the flap side edge. The vortex path and strength are determined, and the process of vortex merging is documented in detail in the mid-chord region. At the higher flap deflection angle, vortex bursting is observed, which may be a factor in the observed higher noise levels for this condition. Results are compared with Reynolds-averaged NavierStokes calculations of a similar configuration (AIAA Paper 98-0768). These measurements and calculations form the basis for instability modeling and unsteady calculations which will ultimately lead to an understanding of flap side-edge noise sources.

Evaluation of PowerFLOW for Aerodynamic Applications

A careful comparison of the performance of a commercially available Lattice-Boltzmann Equation solver (PowerFLOW) was made with a conventional, block-structured computational fluid-dynamics code (CFL3D) for the flow over a two-dimensional NACA-0012 airfoil. The results suggest that the version of PowerFLOW used in the investigation produced solutions with large errors in the computed flow field; these errors are attributed to inadequate resolution of the boundary layer for reasons related to grid resolution and primitive turbulence modeling. The requirement of square grid cells in the PowerFLOW calculations limited the number of points that could be used to span the boundary layer on the wing and still keep the computation size small enough to fit on the available computers. Although not discussed in detail, disappointing results were also obtained with PowerFLOW for a cavity flow and for the flow around a generic helicopter configuration.

Stability Analysis for Noise-Source Modeling of a Part-Span Flap

The relevant and important flow features of the local mean flowfleld In the vicinity of a flap side edge are extracted from an extensive computational and experimental database. Based on the local flow features, possible mechanisms for flow-induced noise sources that are responsible for sound generation at flap side edges are conjectured. Relatively simple flow models for the dominant noise sources are proposed and developed. The models are based on the most amplified instability modes of the local shear flow. Stability analysis of the modeled flowfleld provides an estimate of the frequencies associated with the local flow unsteadiness. Computed frequencies show the proper trend and are in good agreement with NASA Langley Research Centers acoustic microphone array and unsteady surface-pressure measurements.

Hybrid acoustic predictions

Abstract Hybrid acoustic prediction combines the use of unsteady computational fluid dynamics for calculating the properties of acoustic sources with some form of Lighthills acoustic analogy to calculate the resulting far-field noise radiation. This approach avoids some of the difficult issues involved in computing both the acoustic sources and the acoustic propagation in a single computation. However, hybrid acoustic prediction still faces numerous uncertainties associated with the computation of the properties of the acoustic sources, and the coupling of the noise-source computation with the acoustic propagation computation. Several of these uncertainties are discussed, including examples where some of the issues have been successfully resolved.

Development of Computational Aeroacoustics Tools for Airframe Noise Calculations

This paper discusses the development of computational aeroacoustics (CAA) tools for airframe noise analysis and prediction. We review recent progress in this topic, but emphasize our vision for the future development of such tools. Our intention is for this vision to drive future CAA research in directions that will accelerate widespread use of CAA for airframe noise applications. We discuss the needs for accuracy, efficiency, and easy interface with other design tools and illustrate how CAA tools may help future aircraft design. We explain what appears to be achievable in a 20-year time frame, and what goals will probably take longer. Important barrier issues include the effects of numerical dispersion and dissipation, the treatment of highly curved, irregular boundary surfaces, and grid generation. Beyond these largely numerical issues, we discuss the role of physics-based modeling, including turbulence modeling in unsteady flow computations and the importance of developing sophisticated techniques for analyzing results of computations. Numerical simulations combined with the acoustic analogy methodology to predict noise are also reviewed. Finally, we discuss how to use recent advances in measurement techniques for CAA tool validation, which is an integral part of future development.