Foluso Ladeinde

Supersonic Jet Noise from Round and Chevron Nozzles: Experimental Studies

High speed exhaust noise reduction continues to be a research challenge for supersonic cruise business jets as well as for current and future tactical military aircraft. Significant noise reduction may be possible from advanced concepts for controlling instability generated large-scale turbulence structures in the jet shear layer, generally accepted to be the source of aft-angle noise. In response to this opportunity, our team is focused on experimental diagnostic studies and unique instability modeling suited for identifying control strategies to reduce large scale structure noise. The current paper benchmarks the jet noise from supersonic nozzles designed to provide the supporting experimental data and validation of the modeling. Laboratory scale jet noise experiments are presented for a Mach number of Mj = 1.5 with stagnation temperature ratios ranging from Tr=0.75 to 2. The baseline configuration is represented by a round converging-diverging (CD) ideal expansion nozzle. A round CD nozzle with chevrons is included as the first of several planned non-circular geometries directed at demonstrating the impact on large scale structure noise and validating noise prediction methods for geometries of future technological interest. Overexpanded and underexpanded conditions were tested on both nozzle configurations. The resulting data base provides an opportunity to benchmark the statistical characteristics of round and chevron nozzle data. The current paper examines far field spectra, directivity patterns, and overall sound pressure level dependence comparing observed characteristics with the fine scale turbulence noise and large-scale turbulence structure noise characteristics identified by Tam. In addition, the paper probes the effect of chevrons on the developing flow field and suppression of screech tones. Measurements are also reported from a far-field narrow aperture phased array system used to map the acoustic source distribution on the jet axis. The dominant source region, situated between the end of the potential core and the sonic point, was found to agree with the peak amplitude location of the jet near field wavepackets measured using a unique near field array. This observation supports the cause-effect link between large-scale turbulence structures in the shear layer and their dominant contribution to aft radiated far field noise.

Advection by polytropic compressible turbulence

Direct numerical simulation (DNS) is used to examine scalar correlation in low Mach number, polytropic, homogeneous, two‐dimensional turbulence (Ms≤0.7) for which the initial conditions, Reynolds, and Mach numbers have been chosen to produce three types of flow suggested by theory: (a) nearly incompressible flow dominated by vorticity, (b) nearly pure acoustic turbulence dominated by compression, and (c) nearly statistical equipartition of vorticity and compressions. Turbulent flows typical of each of these cases have been generated and a passive scalar field imbedded in them. The results show that a finite‐difference based computer program is capable of producing results that are in reasonable agreement with pseudospectral calculations. Scalar correlations have been calculated from the DNS results and the relative magnitudes of terms in low‐order scalar moment equations determined. It is shown that the scalar equation terms with explicit compressibility are negligible on a long time‐averaged basis. A phy...

Turbulence spectra characteristics of high order schemes for direct and large eddy simulation

Abstract The comparative resolution of the high wavenumber portion of compressible turbulence energy spectrum by some high order numerical schemes is presented in this paper. Included in this are the essentially nonoscillatory (ENO) schemes, the weighted essentially nonoscillatory schemes (WENO), and the compact differencing schemes. The governing equations are the Navier–Stokes equations and the objective is to identify the numerical scheme that best represents the physics of compressible turbulence. Mach numbers M 1 values of 0.1, 0.5 and 0.7 are studied. The compact differencing schemes need filters for numerical stability. It is found in this work that a parameter in the filter scheme provides some flexibility for controlling the physical turbulence energy transfer rate at high wavenumbers, vis-a-vis the numerical dissipation at those scales. Although the ENO schemes do not require filters for numerical stability, the present study shows that the addition of filters improves the energy transfer process at high wavenumbers. Without filtering, with relatively coarse grids, numerical turbulence caused by stencil adaptation persists. This limits the useful wavenumber resolution range of the ENO schemes. The WENO schemes do not require the stabilizing filters but the results tend to be slightly more dissipative. Finally, at low Mach numbers, the current compact differencing and filter scheme formulation gives better results but as the Mach number increases the relative suitability of the ENO method increases.

Supersonic flux-split procedure for second moments of turbulence

A finite volume procedure which combines flux-difference splitting and flux-vector splitting is presented in this paper. The diffusive fluxes are treated explicitly to preserve the upwinding structure of the split Euler fluxes. This procedure is extended in this paper from its essentially laminar or eddy viscosity form to include the equations for the six components of the Reynolds stress tensor and an additional equation for the solenoidal dissipation. The models used for the unclosed terms are presented as are the extensions of the numerical procedures to cover the turbulence equations. To validate the proposed procedures, a compressible turbulent flow over a flat plate at Mach number M∞ = 2.87 and Reynolds number per unit length Re/m of 6.5X10 7 was calculated and compared with experimental measurements. Also simulated was the case when the foregoing boundary-layer flow was made to pass over a ramp. A strong adverse pressure gradient case in which boundary-layer thickness-to-curvature ratio (δ/R c ) is 0.1 was considered. In this case, the supersonic turbulent boundary layer experiences the combined effects of an adverse pressure gradient, bulk compression, and concave streamline curvature.

Advanced Computational-Fluid-Dynamics Techniques for Scramjet Combustion Simulation

The simulation of supersonicmixing and combustion,with a focus on the supersonic combustion ramjet (scramjet), poses many challenges. Even for the nonreacting limit of the problem, the way to properly treat compressibility effects on turbulence when the Mach number is sufficiently high is far from being resolved, independent of previous studies that have advocated the decomposition of the turbulence dissipation of kinetic energy into dilatational and solenoidal components. Complex geometries posed additional difficulties. When combustion dynamics are added to the problem, uncertainty exists in the turbulence–chemistry interaction, shockwave–chemistry interaction, and the general lack of adequate knowledge of the effects of supersonic conditions on turbulence, reaction rates, and flame regimes. Of the three common approaches for modeling turbulence [direct numerical simulation (DNS), large-eddy simulation (LES), and Reynolds-averaged Navier–Stokes equations (RANS)], it is well known that the RANS approach is themost computationally efficient and has the chance of completely modeling realistic aerospace systems. This is followed by LES, whereas DNS is still too costly for realistic engineering problems. However, the success of the RANS approach is problem-dependent, and the procedure needs to be calibrated for every class of problem, making it nonuniversal. Moreover, the approach is inherently steady and cannot deal with unsteady large-scale structures that determine the dynamics ofmany important flow problems. The main issue with LES, relative to RANS, is the computational cost. The special section on advanced simulation of scramjet combustion and mixing in this issue of the AIAA Journal was motivated by the need to assess the current state of the art in this important technological area. The genesis of the efforts originates from the Fluid Dynamics Technical Committee of AIAA, in which a Discussion Group was formed to address the issue. This was followed by a highly successful invited session on the topic at 2009 AIAA Aerospace Sciences Meeting (ASM) in Orlando, Florida. Eight experts on high-speed-combustion modeling were invited to present their work and provide an assessment of the state of the art of scramjet combustion andmixing. TheAIAA JournalEditorial Board agreed to allow the special section on the subject based on its evaluation of the significance. The presenters at the ASM meeting were invited to contribute, and additional international contributors were also brought in to participate. The various contributions went through the normal AIAA Journal review process, the outcome of which is the five papers in this special section of this issue. The first paper is by Ingenito and Bruno, who address some fundamental aspects of supersonic combustion ramjet, with a focus on the physics, as a way to reverse what the authors see as an undue emphasis on numerics over physics. The paper provides a fairly detailed discussion of the effects of compressibility (Mach number) on turbulence, reaction rate, and flame regime. Helicity (the projection of the vorticity vector onto the velocity vector), or streamwise vorticity, is discussed as a way of characterizing the effects of compressibility on turbulent supersonic internal flows. The authors remind the reader that, unlike in subsonic flows, vorticity transport is not exclusively driven by vortex stretching, but also by compressibility and baroclinic effects. They report higher growth rates in high-Mach-number, fully-three-dimensional shear layers compared with incompressible flows, and they consider this to be a source of enhanced mixing at the molecular level, potentially leading to short flames and efficient combustion. One important result from their work is the derivation, on simple dimensional grounds (as in the original 5=3 law), of the decay of turbulence energy with wave number to the power of 8=3, and the consequent observations that the dissipative eddies in supersonic flows are larger than those in subsonic flows. Experiments have supported this observation. Thus, in supersonic flows, the smallest eddiesmight only be able towrinkle flames, without thickening them. Concerning the effects of compressibility on reaction rate, the authors show that compressibility effects on collision frequency could become very important in supersonic flows and that the reaction rate is enhanced by high speed: an effect that should be taken into account when modeling chemical kinetics for supersonic combustion. Compressibility effects are shown to alter the boundaries of the flame regimes relative toBorghi’s diagramorKilmov-William’s. The authors suggest that large Mach numbers raise the possibility of flamelets in eddies. The simulation of hydrogen/air combustion in NASA Langley Research Center experiments under the SCHOLAR program is used to demonstrate some of the theoretical results. The second paper is by Génin andMenon, who address a solution to two of themajor problems that confront high-speed computational fluid dynamics (CFD): 1) the development of schemes that maintain their integrity (accuracy, robustness, etc.) in regions of high discontinuities, as well as in the smooth part of the flow, and 2) the development of versatile turbulence closures that are valid for a wide range of aerospace engineering applications. Génin and Menon address these issues from the perspective of LES, with a focus on adaptable numerical algorithms (via the Harten–Lax–van Leer contact and Harten–Lax–van Leer–Einfeldt hybrid), and dynamically obtain turbulence parameters and subgrid-scale closures that do not have adjustable constants. These approaches are particularly welcome for high-speed flows, in which the development of LES closures is more limited, and the issues of dilatational and solenoidal contribution to dissipation has not been fully resolved. In Génin and Menon’s paper, the compressible part of the turbulent field (or turbulent Mach number) is considered to be small in the high-speed boundary and shear layers, with possible significant contributions in situations in which shock waves interact with boundary or shear layers, as in scramjet combustion flows. The application of the numerical work was based on the facility used in experimental studies at the DLR, German Aerospace Center, which also produced the validation data for the simulations. For the high-speed-combustion problem, a large number of subgrid-scale terms arise that require modeling: subgrid kinetic energy, subgrid gas constant-temperature correlation, subgrid stress, subgrid viscous work, subgrid enthalpy flux, subgrid diffusion of energy, subgrid species flux, subgrid species diffusion–velocity correlation, subgrid mixture-gas constant-temperature correlation, and subgrid mass-fraction–internal-energy correlation. The subgridscale kinetic energy, obtainable via a transport equation, requires that additional terms (triple-velocity correlations, production, dissipation, and pressure dilatation) be modeled. Génin andMenon provide models for the foregoing subgrid-scale terms, except the last four, which are neglected.Note that DNS results in realistic geometries are not yet available for these terms, and so they represent part of the unresolved problem in high-speed-combustion flow simulation. Agreement with experimental data is reported to be fairly good, with discrepancies that are probably due to the use of velocity slip conditions at the walls in the transverse direction. The closure of the AIAA JOURNAL Vol. 48, No. 3, March 2010

Decomposition of High Speed Jet Noise: Source Characteristics and Propagation Effects

Current research programs directed at supersonic engine exhaust noise reduction are demonstrating benefits of 3-4 dBA using passive methods to increase jet mixing and break up shock cells in over-expanded flows. While progress is being made, high speed jet noise continues to be a research challenge for small business jets and tactical military aircraft. The current work benchmarks high speed jet noise using laboratory scale jets for the purpose of a) identifying source and propagation mechanisms, and b) providing validation data for simulation/modeling methods. Laboratory scale experiments are presented over a Mach number range of M = 0.68 to 1.5 with static temperature ratio ranging from Tr = 0.68 to 2. A unique near field rotating phased microphone array technique was used to identify the large-scale turbulence structure noise source and Mach waves in supersonic shock-free jets. A companion paper documents the near field pressure statistics and projection of the convected wave packet to the far field. Validation against the directly measured far field levels quantitatively establishes the large scale structure noise contributions. The combined studies underpin a long term effort to develop modeling methods and new concepts for jet noise suppression based on controlling the evolution of the large-scale turbulence structures.

The First High-Order CFD Simulation of Aircraft: Challenges and Opportunities

The textbook advantages of high-order differencing schemes in computational fluid dynamics (CFD) are well documented. They have also been demonstrated in the literature, albeit for canonical problems. The objective of the work described in this paper is to provide a robust implementation of high-order schemes to permit high fidelity and routine simulation of realistic aerospace systems which usually involve very complex geometries and flow behaviors. The two high-order schemes we have implemented are the compact and weighted essentially non-oscillatory (WENO) schemes. Some challenges were encountered in our efforts to accomplish the foregoing objective. Detailed illustrations of the various challenges are provided, as are some remedies that we have proposed and successfully implemented. This then allows us to illustrate some of the potential advantages of high-order methods for the simulation of realistic aerospace applications. The roles that the authors envision for high-order methods in CFD simulation of realistic aerospace systems are also discussed.

Level-Set Flamelet/Large-Eddy Simulation of a Premixed Augmentor Flame Holder

The objective of the current study is to combine a high-fidelity large eddy simulation (LES) flow solver with a level-set flamelet algorithm for the prediction of premixed turbulent combustion. The same level of high accuracy is implemented for simulation at all speeds. The goal of this work is to accurately predict the unsteady turbulence-flame interaction for realistic industrial combustors with complex geometries. The numerical issues related to the numerical implementation of the LES equations, flamelet model and level-set algorithm are presented in detail. The accuracy of the numerical implementation is verified through comparisons with experimental data for an augmentor flame holder and a turbulent Bunsen burner flame.

Thermodynamic behavior in decaying, compressible turbulence with initially dominant temperature fluctuations

Direct numerical simulation of decaying, isotropic, compressible turbulence in three dimensions is used to examine the behavior of fluctuations in density, temperature and pressure when the initial conditions include temperature fluctuations larger than pressure fluctuations. The numerical procedure is described elsewhere, the initial turbulent Mach number is subsonic, 0.3 to 0.7, and the initial compressible turbulence is characterized as being in one of three states in which the ratios of initial kinetic energy in the compressible modes to total kinetic energy are, respectively, very small, moderate or nearly unity. Only at the lowest values of initial turbulent Mach number and energy ratio do thermodynamic scalings follow the predictions in the literature. For turbulent Mach numbers above 0.3, or for finite values of the kinetic energy ratio, the scalings are more complex. A relationship between turbulent Mach number, compressible pressure and energy ratio, which has been proposed previously for isothe...

Comparative advantages of high-order schemes for subsonic, transonic, and supersonic flows

This computational study aims to verify the order of accuracy of the COMPACT and weighted essentially non-oscillatory (WENO) finite difference schemes implemented in the AEROFLO software, and to identify the comparative advantages of these schemes relative to the low-order (MUSCL-based) schemes for a range of flow problems. The method of manufactured solutions was used to determine the order of accuracy of the spatial differencing schemes. The theoretical sixth-order of accuracy is verified for the COMPACT scheme for subsonic flows, while the theoretical fith-order WENO scheme exhibited a 3.5 order of accuracy for supersonic flows. The MUSCL scheme shows the theoretical secondorder accuracy for all flow regimes. The accuracy results were observed for both Cartesian and curvilinear grids. Several subsonic, transonic, and supersonic calculations were then used to evaluate the results from the highand low-order schemes. For the subsonic and transonic flow configurations, the high-order schemes generally require smaller CPU times, due to their ability to use larger time step sizes or their ability to generate better results with coarser grids as compared to the low-order schemes. For the supersonic flow configurations, both the highand the low-order schemes capture the shock locations very accurately, although the low-order schemes tend to exhibit significantly larger numerical noise in the regions behind the shocks.