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A test campaign was conducted in the open-jet test section of the NASA Langley Research Center's 14- by 22-Foot Subsonic Tunnel on a 10%-scale semispan version of the High-Lift Common Research Model incorporating a leading-edge slat, a trailing-edge flap, and removable high-fidelity main landing gear. Computational simulations were used to support the model development, provide important information such as load distributions along the high-lift elements, and aid in the design of noise reduction devices. Aerodynamic and far-field acoustic measurements were obtained on baseline and acoustically treated model configurations using an ensemble of time-averaged and unsteady surface pressure sensors paired with a traversing 97-microphone phased array viewing the pressure side of the airframe. A deconvolution method incorporating corrections for shear layer acoustic wave decorrelation was employed to determine the locations and strengths of relevant noise sources along the span of the slat and in the vicinity of the flap edges and landing gear. The main goal of the test campaign was to use the surface pressure and array data to evaluate the effectiveness of various slat noise reduction concepts. It was found that slat-gap fillers produced significant broadband noise reduction with minimal aerodynamic performance penalties for all Mach numbers and airfoil angles of attack that were investigated. The test campaign clearly demonstrated the noise reduction benefits that can be obtained by applying appropriate treatments to leading-edge slats on commercial transport-class aircraft. [ABSTRACT FROM AUTHOR]

This paper considers potential sources of error when using the Ffowcs Williams-Hawkings equation to make predictions of airframe noise, which entails a relatively low-speed, uniform incoming flow encountering geometry of varying complexity. Numerical simulations are used to investigate several model problems where Ffowcs Williams-Hawkings integration surfaces are placed on solid surfaces as well as in the flow. Comparisons with the pressure obtained directly from the simulations reveal that when solid surfaces are used, the acoustic calculations can produce erroneous results in upstream directions and when scattering bodies block the line of sight from observers to the source. Using solid surface input data implies ignoring all volumetric source effects, which include noise generation as well as flow effects. Nonuniform flow alone, such as is found in a steady boundary layer, was not found to be a significant source of error, so the amplitude and phase changes induced by turbulent eddies in massively separated flow regions is speculated to be the primary cause of the error. [ABSTRACT FROM AUTHOR]

The current effort characterizes the details of flow interactions and wake interference effects between two cylinders in a tandem configuration. This setup is representative of several component-level flow interactions that occur when air flows over the main landing gear of aircraft. Such interactions are likely to have a significant impact on the noise radiation associated with the undercarriage. This paper focuses on two-dimensional, time-accurate flow simulations of two distinct tandem cylinder flow regimes, associated with short and intermediate separation distances between the two cylinders. Unsteady Reynolds averaged Navier-Stokes simulations using a two-equation turbulence model run at a Reynolds number of 1.66 x [10.sup.5] and a Mach number of 0.166 are presented. Emphasis is placed on understanding both the time-averaged and unsteady flow features between the two cylinders and in the wake of the rear cylinder. Predicted mean-flow quantities and vortex shedding frequencies show reasonable agreement with measured data for both cylinder spacings. Computations for the short separation distance exhibit a nonphysical decay of flow unsteadiness with time: however, the predicted sensitivity of the mean lift coefficient to small variations in the upstream flow angularity explains the asymmetric flowfield observed in the present and previous measurements.

Results demonstrate that high-amplitude, high-frequency acoustic signal, similar to those obtained in high-lift system, is caused by vortex shedding from the trailing-edge slat. The acoustic analysis is based on the Ffowcs Williams and Hawkings equation.

The steady flow calculations over a finite wing were performed using the high-order spatial operators of a computational aeroacoustics (CAA) scheme, as well as a time-marching algorithm. The equations of motion in three dimensions are provided, and there is emphasis on three-dimensional results for low-speed flow over a NACA 0012 wing. Vortex growth was larger than in experimental measurements, although other parameters were well predicted.

A research was conducted to determine the nature of broadband blade noise and determine the reliability of utilizing computational aeroacoustics procedures and parallel computers to further understand broadband blade noise. The two dimensional Navier-Stokes code was utilized in calculating numerical data. Results revealed that computational aeroacoustics techniques are reliable when applied in more realistic problems and that multigrid procedures and parallel computers can be used to obtain realistic mean flows.

High-order and high-bandwidth algorithms, which can function efficiently even in the absence of fine grids, provide accurate solutions for computational aeroacoustics problems. Design of high-bandwidth operators needs to be performed in an appropriate manner to mitigate possible errors in computations involving longer wavelength.

An 18% scale semispan model is used as a platform for examining the efficacy of microphone array processing using synthetic data from numerical simulations. Two hybrid Reynolds-Averaged-Navier-Stokes/Large-Eddy-Simulation (RANS/LES) codes coupled with Ffowcs Williams-Hawkings solvers are used to calculate 97 microphone signals at the locations of an array employed in the NASA Langley Research Center 14×22 tunnel. Conventional, DAMAS, and CLEAN-SC array processing is applied in an identical fashion to the experimental and computational results for three different configurations involving deploying and retracting the main landing gear and a part-span flap. Despite the short time records of the numerical signals, the beamform maps are able to isolate the noise sources, and the appearance of the DAMAS synthetic array maps is generally better than those from the experimental data. The experimental CLEAN-SC maps are similar in quality to those from the simulations indicating that CLEAN-SC may have less sensitivity to background noise. The spectrum obtained from DAMAS processing of synthetic array data is nearly identical to the spectrum of the center microphone of the array, indicating that for this problem array processing of synthetic data does not improve spectral comparisons with experiment. However, the beamform maps do provide an additional means of comparison that can reveal differences that cannot be ascertained from spectra alone. [ABSTRACT FROM AUTHOR]

The present article concerns a commonly used methodology for the numerical simulation of acoustic emission and propagation phenomena. We consider the so-called multi-stage hybrid acoustic approach, in which a given noise problem is simulated via a sequence of weakly coupled computations of noise generation and acoustic propagation stages, wherein the simulation of the propagation stage is based on advanced Computational AeroAcoustics (CAA) techniques. The paper introduces an original forcing technique, namely, the Non-Reflective Interface (NRI), to enable the transfer of an acoustic signal from an a priori noise generation stage into a CAA-based acoustic propagation phase. Unlike most existing forcing techniques, the NRI is non-reflective (or anechoic) in nature and, therefore, can properly handle the backscattering effects arising during the noise propagation stage. This attribute makes the NRI-based weak-coupling procedure and the associated CAA-based hybrid approach compatible with a larger variety of realistic noise problems (such as those involving installed configurations in wind tunnel experiments, for instance). The NRI technique is first validated via several test cases of increasing complexity and is then applied to two aerodynamic noise problems. Copyright © 2015 John Wiley & Sons, Ltd. Copyright © 2015 John Wiley & Sons, Ltd. [ABSTRACT FROM AUTHOR]

Abstract: Assessing solution error continues to be a formidable task when numerically solving practical flow problems. Currently, grid refinement is the primary method used for error assessment. The minimum grid spacing requirements to achieve design order accuracy for a structured-grid scheme are determined for several simple examples using truncation error evaluations on a sequence of meshes. For certain methods and classes of problems, obtaining design order may not be sufficient to guarantee low error. Furthermore, some schemes can require much finer meshes to obtain design order than would be needed to reduce the error to acceptable levels. Results are then presented from realistic problems that further demonstrate the challenges associated with using grid refinement studies to assess solution accuracy. [ABSTRACT FROM AUTHOR]

The relevance of geometric details to the generation and propagation of noise from leading-edge slats is considered. Typically, such details are omitted in computational simulations and modelscale experiments thereby creating ambiguities in comparisons with acoustic results from flight tests. The current study uses two-dimensional, computational simulations in conjunction with a Ffowcs Williams-Hawkings (FW-H) solver to investigate the effects of previously neglected slat "bulb" and "blade" seals on the local flow field and the associated acoustic radiation. The computations show that the presence of the "blade" seal at the cusp in the simulated geometry significantly changes the slat cove flow dynamics, reduces the amplitudes of the radiated sound, and to a lesser extent, alters the directivity beneath the airfoil. Furthermore, the computations suggest that a modest extension of the baseline "blade" seal further enhances the suppression of slat noise. As a side issue, the utility and equivalence of FW-H methodology for calculating farfield noise as opposed to a more direct approach is examined and demonstrated. [ABSTRACT FROM AUTHOR]

Abstract: Assessing solution error continues to be a formidable task when numerically solving practical flow problems. Currently, grid refinement is the primary method used for error assessment. The minimum grid spacing requirements to achieve design order accuracy for a structured-grid scheme are determined for several simple examples using truncation error evaluations on a sequence of meshes. For certain methods and classes of problems, obtaining design order may not be sufficient to guarantee low error. Furthermore, some schemes can require much finer meshes to obtain design order than would be needed to reduce the error to acceptable levels. Results are then presented from realistic problems that further demonstrate the challenges associated with using grid refinement studies to assess solution accuracy. [Copyright &y& Elsevier]

The long-term objective of the research described in this paper is to use CAA methodology and parallel computers to increase the understanding of broadband blade noise. In a systematic progression towards simulations of completely realistic configurations and conditions, some simplified problems that address the important features of the flow are investigated. A two-dimensional, Euler code, implemented using the message passing library and Fortran 90 on the IBM SP2, is used to perform the calculations. Results are presented for the interaction of a vortical gust and a flat plate. [ABSTRACT FROM AUTHOR]

The slat of a high-lift wing can be a significant noise source during approach and landing. This paper describes an experimental study of the effects of two passive devices (namely, slat extensions and a cove filler) on the aerodynamics and acoustic radiation from a 30P30N model. Test conditions include three effective free-air angles of attack of 5.5, 7.5, and 13 deg, as well as three Reynolds numbers of 1.2, 1.5, and 1.71×106. Steady surface pressure measurements are used to assess the aerodynamics near the midspan section. Unsteady surface pressure and far-field acoustic array measurements are performed to evaluate the near- and far-field pressure fluctuations, respectively. Delay-and-sum beamforming is applied to localize the noise sources and to provide integrated spectra. The slat cusp extensions barely affect the mean Cp distribution, whereas the longest extension achieves up to a 5.4 dB reduction in the overall sound pressure level. The slat-cove filler yields an even higher noise reduction that is lower than the baseline by up to 14 dB. Even though the aerodynamic assessment of the cove filler is incomplete due to the covered pressure ports, the mean Cp distribution over the remaining surface of the model indicates a rather small change compared with the baseline measurements. [ABSTRACT FROM AUTHOR]

The work presented in this paper is a culmination of a multiyear joint experimental and computational effort to explore the feasibility of using active flow control (AFC) on a simple hinged flap system for recovering lift comparable to a conventional high-lift system consisting of Fowler flaps. The baseline configuration chosen for this work is the high-lift version of the NASA Common Research Model (CRM), which is a representative modern aircraft consisting of wing, fuselage, nacelle/pylon, slats, Fowler flaps, and slat and flap brackets. A simplified high-lift (SHL) system was created by replacing the Fowler flaps and flap brackets with a simple hinged flap system equipped with integrated modular AFC cartridges on the suction surface of the flap shoulder, and the resulting geometry is known as the CRM-SHL-AFC configuration. Parametric studies were conducted in the earlier phases of this effort to numerically evaluate and downselect the more efficient AFC designs for wind-tunnel tests. These simulations were performed with the PowerFLOW® code, which is a lattice-Boltzmann-based computational fluid dynamics code. Good agreement was reported in a previous paper between the numerical results and the experimental data for the lift characteristics of the CRM-SHL-AFC configuration as a function of actuation levels at the nominal landing conditions. The current effort is focused on demonstrating the applicability of the PowerFLOW code for predicting the aerodynamic performance of the conventional and the AFC-enabled simplified high-lift CRM configurations for a broad angle of attack range, including maximum lift conditions. [ABSTRACT FROM AUTHOR]