Jan Delfs

An overlapped grid technique for high resolution CAA schemes for complex geometries

A high order/high resolution grid communication algorithm was developed for use in combination with the high resolution DRP scheme of Tarn and Webb. In order to establish this grid communication a special high order, though straight forward interpolation scheme is employed at the interfaces between two curvilinear grids of arbitrary shape. The interpolation procedure is local and shows excellent spectral characteristics such that for a given resolution the application range of the DRP-scheme is not narrowed. Twodimensional tests have been done of which may be concluded that the method is robust since additional time step limitations or numerical instabilities have not been observed. A reliable grid overlap detector algorithm for high order metric representation was developed and tested, which is capable of identifying points inside and outside arbitrarily curved domains and subdomains. Geometric information is needed only on the surface of the domains such that their interior does not need to be meshed. Benchmark examples are presented to show the validity of the approach.

Spectral broadening of jet engine turbine tones

The process of turbulent scattering is studied for the generic experiment conducted by Candel et al.1 applying an analytic weak scattering model and CAA computations. For the analytic weak scattering model an approximate form of the Lilley equation is used. The source terms of this equation are in terms of the turbulence and the incident acoustic field. In the CAA simulations the wave equation proposed by Pierce for sound in fluids with unsteady inhomogeneous flow is integrated. The unsteady turbulent base-flow is modeled using a stochastic method to generate turbulence with locally varying turbulence features as provided by time-averaged RANS. To study the spectral broadening effect analytically and computationally, the experimental set-up of Candel is considered, which involves an omnidirectional sound source, located on the axis of a round jet. The analytical predictions show very good agreement with the general trends as measured by Candel for an observer position normal to the jet axis. The computations reveal a spectral shape, which is in good agreement with those found in the experiments.

Specifcation of Porous Materials for Low-Noise Trailing-Edge Applications

Systematic testing of the microstructural and aeroacoustic properties of porous metals applicable as low-noise trailing-edge (TE) treatments has been initiated within the Col- laborative Research Center SFB 880|Fundamentals of High-Lift for Future Civil Aircraft. Generic TE noise experiments were performed at Re = 0.8x10^6 to 1.2x10^6 in DLRs open-jet AWB facility. Complementary flow measurements in the closed test section MUB wind-tunnel of the TU Braunschweig served to quantify the induced aerodynamic effects. The presented database forms part of an ongoing cumulative effort, combining experimental and numerical methods, to gain a deeper understanding of the prevalent TE noise reduction mechanisms. For the large variety of porous materials tested herein a clear dependence of the achieved broadband noise reduction (reaching 2-6 dB at maximum) on the flow resistivity was identified. Basic design recommendations for material resistivity and pore sizes, the latter to minimize high-frequency self-noise contributions, were deduced for low-noise TE applications. An acoustic nearfield pressure release across the porous region, adversely coupled with a loss in lift performance for porous TE replacements, appears as the major noise-reduction requirement.

Enhanced Capabilities of the Aeroacoustic Wind Tunnel Braunschweig

The Aeroacoustic Wind Tunnel Braunschweig is DLR’s small, high-quality test facility for aero-acoustic noise measurements. After years of mainly pure acoustic measurements the actual and the future research work will focus on the combined acquisition of aerodynamic and acoustic data. In order to prepare the AWB for the next decades a modernization of this wind tunnel was initiated targeting the enhancement of the aerodynamic properties while the excellent acoustic properties should at least be preserved. Based on the assessment of AWB’s original acoustic and aerodynamic characteristics a new acoustic wall-treatment in combination with new sound absorbing turning vanes was installed in the flow circuit. In order to prevent the highly bended downwash flow caused by airfoils tested at high angles of attack from impinging on the test section floor a new collector was installed that can be moved upstream and vertically down. Finally the driven efforts resulted in an increase of the maximum flow velocity of about 8% while first the background noise levels for an empty test section were preserved and second the use of the adaptive collector lead to a significant background noise reduction for airfoil tests.

A CAA Based Approach to Tone Haystacking

The far-field noise spectra of jet engines show for certain jet configurations and turbine tones a characteristic spectral broadening effect, causing a reduction of tone peaks in favor of a more distributed spectral hump around each tone frequency. This haystacking effect likely occurs due to the interaction of the turbine tones with the unsteady turbulent jet shear layer. A better understanding of this effect may help to utilize it for noise reduction purposes. Furthermore, the effect is of interest for the measurement of tone sources in an open acoustic wind tunnel test section, since the tone will be scattered in the open jet shear layer. A correction for this measured broadening effect is desirable. A non-empirical computational approach to predict tone haystacking as a function of Reynolds number/jet shear layer characteristics is currently missing. This paper reports about ongoing work to utilize Computational Aeroacoustics (CAA) methods for the prediction of haystacking. In a first step CAA techniques are applied to simulate the propagation of tones through the time averaged steady exhaust of a jet engine. To simulate the haystacking effect with CAA, the unsteady turbulent base-flow is modeled with a 4D synthetic turbulence method. The employed RPM approach generates turbulence with all local statistical features as predicted by time-averaged RANS. To study the spectral broadening effect computationally, the experimental set-up of Candel is considered first, which involves an omnidirectional sound source located on the axis of a round jet. The analytical predictions show very good agreement with the general trends as measured by Candel for an observer position normal to the jet axis. The computations reveal a spectral shape, which is in good agreement with those found in the experiments. In a next step the methodology is combined with the exhaust problem to simulate sound propagation through the unsteady turbulent exhaust.

Sound Generation from Gust-Airfoil Interaction Using CAA-Chimera Method

The sound field around a finite-span airfoil subjected to a two-dimensional periodic gust is simulated using CAA-Chimera method. The perturbation equations about the mean flow are used as governing equations in solving this problem. The Chimera methodology in conjunction with the high-order central difference DRP (Dispersion-Relation-Preserving) scheme, so called CAA-Chimera, has been employed to show the usefulness of the approach for realistic aeroacoustics problems. The body fitted grid is only used in a very narrow zone close to the curved body. The primary computational grid is a Cartesian grid. Special non-homogeneous inflow boundary conditions are derived to generate incoming disturbances with different angles of incidence. A multi-domain Cartesian grid system is used in which local resolution enhancement is achieved by stepping the grid size of neighboring Cartesian grid blocks by factors of two. In the area of two Cartesian grid interfaces, a combination of DRP and a specially optimized cell-centered high order differencing scheme is implemented not requiring any interpolations. In a systematic study the influence of different gust wave numbers and the airfoil camber on the radiated noise is described. The effect of mean flow compressibility on noise radiation is also studied.

A Testbed for large scale and high Reynolds number Airframe Noise Research

Airframe noise testing in open jet wind tunnels is and will be an indispensable means to (i) identify airframe noise sources and develop noise reduction technologies and (ii) to provide validation data in order to support the development of numerical acoustic methods like e.g. CAA codes. Usually such wind tunnel experiments were conducted in small research facilities like DLRs AWB (nozzle size 1.2 m x 0.8 m) using 2-dimensional wind tunnel models featuring a retracted chord length of about 300 to 400 mm. In very rare cases full aircraft models of 1:7.5 to 1:11 scale were tested in the Large Low-Speed Facility of DNW (DNW-LLF), which provides a nozzle size of 8 x 6 m2 and a maximum flow speed of 78 m/s. Even though these full aircraft models provide almost realistic 3-dimensional flow conditions the mean retracted chord length is pretty similar to the 2D test cases. Therefore most airframe noise tests are restricted to a Reynolds number regime between Re=1.0*106 and Re=2.0*106. This finally means that the aerodynamic conditions during most airframe noise wind tunnel experiments do not comply with realistic operating conditions. In order to overcome these deficiencies DLR, Airbus and EADS-IW decided to design a new large scale test bed for airframe noise testing. This plan could be realized in the mainframe of the German national founded research project FTEG (Flight Physics Technologies for Green Aircraft).The finally designed high lift system features a retracted chord length of 1200 mm and a wing span of 7200 mm. It was tested for representative approach conditions and Reynolds numbers up to 5.0*106. By means of this wind tunnel model high Reynolds number slat noise data were acquired that later serve for the validation of the DLR CAA code PIANO. It reveals the slat noise spectra did not contain any relevant tonal slat noise components that are well known from small scale experiments. Dedicated Reynolds number variations gave evidence that (i) low Reynolds number tests on high lift airfoils do not completely represent the broadband sound radiation typical for respective full scale components and that (ii) the tonal components are related to more than one generation mechanism. The new large scale F15LS high lift system successfully proved its ability to enable the assessment of noise reduction technology and the acquisition of validation data for numerical methods under almost realistic flow conditions.

Handling of Non-Periodic Contra Rotating Open Rotor Data

In this contribution, we present two different concepts to handle the non-periodic nature of Contra Rotating Open Rotor (CROR) if the front and the aft rotor rotate with a slightly different rotational speed. The first procedure that is presented consists of a correction matrix applied to the source data used in the DLR FWH-code APSIM+. For periodic data the correction matrix reduces to the identity matrix, thereby recovering the standard Fourier transformation. The second method is based on the Vanicek approximation, and consists of a successive least-square approximation of non-periodic data. The developed methods are tested with artificially generated data, illustrating the ability to accurately representing non-periodic data. A comparison between the two methods shows that the first method is more accurate than the Vanicek approximation. Preliminary results on actual non-periodic CROR data reveals the influence of the non-periodic correction as compared to uncorrected data, i.e., differences ranging up to 10 dB are seen for the considered cases.

Simulation of Sound Generation by Vortices passing the Trailing Edge of Airfoils

The interaction of vorticity with the trailing edge of airfoils is an important source of airframe noise. In order to study this process, in the present paper the interaction of a single vortex with the trailing edge of a symmetric Joukowsky airfoil under 0° angle of attack is calculated by the solution of nonlinear disturbance equations with a RANS mean flow in the Mach number range from 0.2 to 0.5. Nonlinearities are taken into account up to first order. The emphasis of the present study rests on the assessment of the influence of the Mach number and non-linearities on the generated sound field. It has been found that the sound intensity downstream of the profile scales approximately like M**4, which is near to the theoretical value. Taking into account non-linearities increases the sound pressure level in case of mathematically negative rotating initial vortices and decreases it in case of positive rotating ones. This should be kept in mind in the assessment of solutions obtained by the linearized Euler equations.

Installation Effects of a Propeller Mounted on a High-Lift Wing with a Coanda Flap. Part I: Aeroacoustic Experiments

In this contribution, we present aeroacoustic experiments concerning installation effects of propellers. Such installation effects are important as they can significantly alter the sound radiation as compared to an isolated propeller. For this purpose, detailed experiments have been conducted in the NWB aeroacoustic wind tunnel in Braunschweig, Germany. The considered geometry is a nine-bladed propeller installed in front of a high-lift wing (employing a Coanda flap). The results illustrate the influence of propeller rotational speed, blade pitch angle, wind tunnel velocity, and angle of attack variations on the sound radiation. Furthermore, with a source localisation technique insight is gained in the dominant sound sources, and reveals the importance of periodic as well as broadband noise for the considered geometry.