Markus Olander Burak

Validation of a Time- and Frequency-Domain Grazing Flow Acoustic Liner Model

In todays aeroengines, acoustic liners are extensively used to suppress noise. To optimize their placement and tuning, there is a need for acoustic liner models capturing their effect. Traditionally, the presence of a mean flow has been accounted for through the Ingard or later the generalization in form of the Myers boundary condition. This paper shows that a direct use of nominal impedance for the acoustic liner is justified if the mean flow is properly accounted for by the flow equations. An accurate assessment of the acoustic liner in the presence of grazing flows can be obtained without using an acoustic liner model accounting for the flow, for example, the Myers boundary condition. Validations have been made for both time and frequency-domain solvers using large-eddy simulations and linearized Navier-Stokes equations.

Micro-jet flow control for noise reduction of a supersonic jet from a practical C-D nozzle

A convergent-divergent nozzle similar to those used on high performance tactical aircraft has been built with micro jets at the trailing edge. These micro jets are used to increase mixing in the shear layer between the core flow and a secondary flow. Near field pressure measurements, far field acoustic measurements, as well as PIV measurements have been taken at a core flow Mach number of 1.56 and at a secondary flow Mach numbers of 0.0, 0.1, and 0.3. These results are compared with a baseline nozzle with the same geometry to see what affect the micro jets have on the jet noise production of the nozzle. A numerical simulation of the micro jets has been done using a large eddy simulation. The numerical results are compared with the experimental results. The micro jets are shown to decrease the sound pressure levels for low frequencies and increase them for high frequencies. These effects become more prominent as the secondary flow of Mach number increases. The presence of the micro jets decreases the OASPL of the core flow for most cases. The azimuth direction that benefits most changes from forward observation angles to aft observation angles as the secondary flow increases.

SUPERSONIC JET NOISE FROM A CONICAL C-D NOZZLE WITH FORWARD FLIGHT EFFECTS

Flow and far-field noise measurements are taken on a conical ConvergentDivergent nozzle similar to the nozzles employed on high-performance tactical jets. Matching flow and far-field computations are presented, produced by Large Eddy Simulation and the Kirchhoff integral method. The conditions examined are those in which the nozzle is operated at its design Mach number of 1.56 while forward flight is simulated at Mach numbers of 0.1, 0.3 and 0.8. Both measurement and LES show that increasing forward flight Mach number to the high subsonic range shortens the initial shock cell size, and weakens the shock cells induced by the nozzle throat relative to the shock cells induced by the nozzle lip. LES shows that high forward flight speed substantially reduces the noise radiated into the forward quadrant where shock noise is dominant. It also removes the screech tone entirely.

Experimental and Numerical Investigation of a Supersonic Convergent-Divergent Nozzle

Experiments and large eddy simulations of a conical convergent divergent nozzle, similar to the nozzles employed on high-performance tactical jets, have been carried out. The nozzle has been studied at a slightly overexpanded condition with a nozzle pressure ratio of 4.0 (a design nozzle pressure ratio 4.1). The primary nozzle is surrounded by a secondary jet operated at three different conditions: Mach 0.1, 0.3, and 0.8. This secondary jet provides a rough forward-flight simulation over a limited axial range. It has been found that the shock pattern in the jet plume is directly influenced by the secondary jet Mach number due to the outer geometric shape of the nozzle. As the flow follows the boat-tail, it locally increases the pressure, making the nozzle operate in a more overexpanded condition. Large eddy simulations of a primary jet with forward flight fully simulated are in very good agreement with the experiments regarding both flow and acoustics. A numerical investigation of screech mode character for the case with the lowest secondary flow Mach number has been made. This shows that the screech mode is mainly built up by two counter-rotating modes with a tangential mode number of I. In the numerical simulations, additional flow-adaptive dissipation has been added in order to stabilize the solution around the shocks. The effect on the flow and acoustics is found to be very small if a shock sensor combined with a damping threshold is used.

Supersonic turbojet noise reduction

Observations and Large Eddy Simulations are presented of a supersonic jet from a nozzle representative of high-performance military aircraft such as the Saab Gripen. The nozzle has a design Mach number of 1.56 and is examined at its design condition with a surrounding secondary flow at Mach numbers of 0.0, 0.1 and 0.3. The nozzle is investigated in its unmodified state and also with the addition of chevrons and microjets. Detailed flow-field velocity measurements of the jets and far-field noise measurements are presented and the noise results are scaled to represent the effects of the chevrons and microjets on airport neighbors. Chevrons and internal fluidic injection by microjets each reduce the noise generated by the main jet. And substantially reduce the noise footprint around the airport. The numerical simulation technique, correctly predicts the flow and noise not only the baseline case, but also the noise reduction by both chevrons and microjects.

Comparison of Flow Control Methods Applied to Conical C-D Nozzles

This paper presents an overview of a joint Experimental/Numerical project sponsored by the Swedish Defense Materiel Administration (FMV) to apply flow control techniques to reduce the noise from high-performance military aircraft such as the Saab Gripen. At University of Cincinnati chevrons and trailing-edge fluidic injection were tested and compared with secondary flow simulating forward flight. At Chalmers University the same conditions were simulated with Large Eddy Simulation and Kirchhoff integral method. Both flow control approaches produced significant reductions in community noise. The area for a given peak perceived noise level was reduced by 20-27% by chevrons and by 23-38% by microjets.

Forward flight effects on the shock structure from a chevron C-D nozzle

Detailed PIV was undertaken on a heated jet from a practical militarystyle convergent-divergent nozzle with a design Mach number of 1.56 operated at its design condition. Three secondary flow Mach numbers were examined, 0.1, 0.3 and 0.8 in addition to measurements with the secondary flow turned off. These measurements were taken with and without chevrons present. Shock structure changes in location, shock angle and shock strength were determined. Chevrons decrease shock strength in an axial plane passing through the valleys between chevrons, but increase shock strength in the axial plane passing through the chevron tips. At the highest secondary flow, the throat shock strength is reduced to a negligible level. In the near-nozzle region chevrons increase the turbulent kinetic energy.

Acoustic Mode Analysis of a Supersonic Jet

The dynamics of a supersonic jet emanating from a high-aspect-ratio rectangular nozzle is investigated using dynamic mode decomposition. In the analysis, dynamic mode decomposition is applied to flowfield data obtained from a large-eddy simulation and the significance of each extracted mode to the snapshot data series is determined using a sparsity-promoting method. Furthermore, in order to assess the validity of the resulting modes, they are compared with the unprocessed large-eddy simulation data in terms of frequency content, sound pressure levels, and space–time correlations. It is found that the accuracy of the modes predicted by the dynamic mode decomposition is sufficient and that the modes deemed as most important give a good representation of the jet flow dynamics. Moreover, the jet flow, at the simulated conditions, is known (from experiments) to be dominated by a screech mechanism. The dynamic mode decomposition methodology, as set up in the present study, is able to isolate the jet screech mode, and the frequency of the extracted mode is in excellent agreement with previous observations.

Flow and noise predictions for a mixer-ejector engine configuration using LES

Large-eddy simulations of the compressible o w and acoustic eld for a mixer-ejector nozzle conguration have been carried out. The conguration consists of two mixing enhancers. Firstly an internal mixer with a lobed trailing edge forcing the core and bypass o ws to mix. Secondly, a section allowing ambient air to enter the ejector and to mix with the core/bypass o w. This two-step mixing arrangement creates three-dimensional structures of widely varying length scales. This puts a high demand on mesh resolution and a robust solution technique in order to achieve high quality simulation results. The o w simulation predicts a highly anisotropic turbulence downstream of the mixers, which indicates that turbulence models based on the Boussinesq assumption are unsuitable for this type of o w. The computational domain is discretized using a block-structured boundary-tted mesh with 219 mesh blocks and approximately 24 10 6 nodes. The choice of a block-structured grid is based on previous experience where this type of mesh has provided good numerical accuracy for both o w and acoustics in combination with the numerical scheme used. However, for a complex geometry such as the mixer-ejector concept, it is a challenge to construct a reasonable block topology for a structured mesh. Nevertheless, this paper shows that this may be done and that a high quality block-structured mesh can be obtained through the use of generalized interfaces with hanging node renemen t/coarsening. The Favre-ltered Navier-Stokes equations were solved using a nite volume method with a low-dissipation third-order upwind-biased scheme for the convective uxes, a secondorder centered dierence approach for the viscous uxes, and a three-stage second-order Runge-Kutta technique in time. A compressible form of Smagorinsky’s subgrid-scale model is used for computation of the subgrid-scale stresses.

Dynamic Mode Decomposition Applied to a Detached-Eddy Simulation of Separated Nozzle Flow

The paper presents results from a Dynamic Mode Decomposition (DMD) of data from a Detached Eddy Simulation (DES) of a separated flow inside a truncated ideal nozzle. Two cases of different pressure ratios were studied. Sparsity-Promoting algorithm along with computed optimal mode amplitudes were used to determine importance of individual modes. An ovalization mode (a mode with azimuthal wavenumber m = 2) was found for the lower pressure ratio and was linked to a peak in spectra from probe data. At the higher pressure ratio a helical mode (azimuthal wavenumber m = 1) was found and linked to a peak in spectra from probe data and the nozzle side-load spectrum. The paper shows the potential for using DMD for separated nozzle flows to identify im- portant periodic flow behavior but also underlines the challenges that the method faces, such as noise from resolved turbulence and difficulty identifying modes within the broad low-frequency-range of the side-load spectrum.