Lennart S. Hultgren

Demonstration of Separation Delay With Glow-Discharge Plasma Actuators

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Predictions of Separated and Transitional Boundary Layers Under Low-Pressure Turbine Airfoil Conditions Using an Intermittency Transport Equation

A new transport equation for the intermittency factor was proposed to predict separated and transitional boundary layers under low-pressure turbine airfoil conditions. The intermittent behavior of the transitional flows is taken into account and incorporated into computations by modifying the eddy viscosity, mu(sub t), with the intermittency factor, gamma. Turbulent quantities are predicted by using Menters two-equation turbulence model (SST). The intermittency factor is obtained from a transport equation model, which not only can reproduce the experimentally observed streamwise variation of the intermittency in the transition zone, but also can provide a realistic cross-stream variation of the intermittency profile. In this paper, the intermittency model is used to predict a recent separated and transitional boundary layer experiment under low pressure turbine airfoil conditions. The experiment provides detailed measurements of velocity, turbulent kinetic energy and intermittency profiles for a number of Reynolds numbers and freestream turbulent intensity conditions and is suitable for validation purposes. Detailed comparisons of computational results with experimental data are presented and good agreements between the experiments and predictions are obtained.

Wave Computation in Compressible Flow Using Space-Time Conservation Element and Solution Element Method

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.

Measurements in Separated and Transitional Boundary Layers Under Low-Pressure Turbine Airfoil Conditions

Detailed velocity measurements were made along a flat plate subject to the same dimensionless pressure gradient as the suction side of a modern low-pressure turbine airfoil. Reynolds numbers based on wetted plate length and nominal exit velocity were varied from 50,000 to 300,000, covering cruise to takeoff conditions. Low and high inlet free-stream turbulence intensities (0.2 and 7 percent) were set using passive grids. The location of boundary-layer separation does not depend strongly on the free-stream turbulence level or Reynolds number, as long as the boundary layer remains nonturbulent prior to separation. Strong acceleration prevents transition on the upstream part of the plate in all cases. Both free-stream turbulence and Reynolds number have strong effects on transition in the adverse pressure gradient region. Under low free-stream turbulence conditions, transition is induced by instability waves in the shear layer of the separation bubble. Reattachment generally occurs at the transition start. At Re =50,000 the separation bubble does not close before the trailing edge of the modeled airfoil. At higher Re, transition moves upstream, and the boundary layer reattaches. With high free-stream turbulence levels, transition appears to occur in a bypass mode, similar to that in attached boundary layers, Transition moves upstream, resulting in shorter separation regions. At Re above 200,000, transition begins before separation. Mean velocity, turbulence, and intermittency profiles are presented.

Near Field Screech Noise Computation for an Underexpanded Supersonic Jet by the CE/SE Method

The space-time conservation element and solution element (CE/SE) method is employed to numerically study the near-field axisymmetric screech-tone noise of a typical underexpanded circular jet issuing from a sonic nozzle. For the computed cases, corresponding to fully expanded Mach numbers of 1.10, 1.15 and 1.19, the self-sustained feedback loop is automatically established. The computed shock-cell structure, acoustic wave length, screech tone frequencies, and sound pressure levels are in good agreement with experimental results.

Noise-Source Separation Using Internal and Far-Field Sensors for a Full-Scale Turbofan Engine

Noise-source separation techniques for the extraction of the sub-dominant combustion noise from the total noise signatures obtained in static-engine tests are described. Three methods are applied to data from a static, full-scale engine test. Both 1/3-octave and narrow-band results are discussed. The results are used to assess the combustion-noise prediction capability of the Aircraft Noise Prediction Program (ANOPP). A new additional phase-angle-based discriminator for the three-signal method is also introduced.

Jet Screech Noise Computation

The near-field screech-tone noise of a typical underexpanded circular jet issuing from a sonic nozzle is simulated numerically. The self-sustained feedback loop is automatically established in the simulation. The computed shockcell structure, acoustic wave length, screech-tone frequencies, and sound pressure levels in the near field are in good agreement with existing experimental results. I. Introduction J ET noise is an important and still challenging topic in aeroacoustics. Many of its aspects are of primary practical importance, and the associated complicated physical phenomena are the topic of many experimental and theoretical investigations. In Refs. 1‐4, a comprehensive discussion and further references are provided. Under/overexpanded supersonic jets emit mixing noise, broadband shock-associated noise, as well as screech tones under certain conditions. The mixing noise is directly associated with large-scale structures, or instability waves, in the jet shear layer, and the broadband shock-associated noise is caused by the interaction of these waves with the shock-cell structure in the jet core. The screech tones arise due to a feedback loop involving the large-scale structures developing in the jet shear layer, their interaction with the jetcore shock-cell structure producing upstream propagating acoustic waves, and regeneration of the large-scale structures at, or in the vicinity of, the nozzle lip. This feedback loop leading to screech tones is sensitive to small changes in the system conditions, and the understanding of the phenomena to date is based mostly on experimental observations. 5−9 Screech is of particular interest not only because of general noise-reduction concerns, but also because of po

Noise Computation of a Shock-Containing Supersonic Axisymmetric Jet by the CE/SE Method

The space-time conservation element solu- tion element (CE/SE) method (l) is em- ployed to numerically study the near-field of a typical under-expanded jet. For the computed case-a circular jet with Mach number Mj = 1.19-the shock-cell struc- ture is in good agreement with experimen- tal results (2, 31. The computed noise field is in general agreement with the experi- ment, although further work is needed to properly close the screech feedback loop.

Direct calculations of waves in fluid flows using high-order compact difference scheme

The solution of the unsteady Euler equations by a sixth-order compact difference scheme combined with a fourth-order Runge-Kutta method is investigated. Closed-form expressnions for the amplification factors and their corresponding dispersion correlations are obtained by Fourier analysis of the fully discretized, two-dimensional Euler equations. The numerical dissipation, dispersion, and anisotropic effects are assened. The CFL limit fontable calculations is about 0.8. For a CFL number equal to 0.6, the smallest wavelength which is resolved without numerical damping is about six-eight grid nodes. For phase speeds corresponding to acoustic waves, the corresponding time period is resolved by about 100-300 time steps

Vortex dynamics simulation in aeroacoustics by the space-time conservation element and solution element method

The space-time conservation element solution element (CEASE) method is employed to numerically study several vortex dynamics problems in aeroacoustics. The scheme possesses low dispersion error and the implementation of non-reflecting boundary conditions is simple and effective. The scheme is advantageous for vorticity computation since the spatial derivatives of the flow variables are computed as an integral part of the solution, i.e. no further numerical differentiation with associated loss of accuracy is needed. The scheme is tested for several nonlinear vortex-shock and vortexblade problems to demonstrate its robustness. The method is found to be valuable for vortex simulation as well as computational aeroacaustics.