William M. Humphreys

Flap Edge Aeroacoustic Measurements and Predictions

An aeroacoustic model test has been conducted to investigate the mechanisms of sound generation on high-lift wing configurations. This paper presents an analysis of flap side-edge noise, which is often the most dominant source. A model of a main element wing section with a half-span flap was tested at low speeds of up to a Mach number of 0.17, corresponding to a wing chord Reynolds number of approximately 1.7 million. Results are presented for flat (or blunt), flanged, and round flap-edge geometries, with and without boundary-layer tripping, deployed at both moderate and high flap angles. The acoustic database is obtained from a small aperture directional array (SADA) of microphones, which was constructed to electronically steer to different regions of the model and to obtain farfield noise spectra and directivity from these regions. The basic flap-edge aerodynamics is established by static surface pressure data, as well as by computational fluid dynamics (CFD) calculations and simplified edge flow analyses. Distributions of unsteady pressure sensors over the flap allow the noise source regions to be defined and quantified via cross-spectral diagnostics using the SADA output. It is found that shear layer instability and related pressure scatter is the primary noise mechanism. For the flat edge flap, two noise prediction methods based on unsteady-surface-pressure measurements are evaluated and compared to measured noise. One is a new causality spectral approach developed here. The other is a new application of an edge-noise scatter prediction method. The good comparisons for both approaches suggest that the prediction models capture much of the physics. Areas of disagreement appear to reveal when the assumed edge noise mechanism does not fully define the noise production. For the different edge conditions, extensive spectra and directivity are presented. The complexity of the directivity results demonstrate the strong role of edge source geometry and frequency in the noise radiation. Significantly, for each edge configuration, the spectra for different flow speeds, flap angles, and surface roughness were successfully scaled by utilizing aerodynamic performance and boundary-layer scaling methods developed herein.

A Deconvolution Approach for the Mapping of Acoustic Sources (DAMAS) Determined from Phased Microphone Arrays

Current processing of acoustic array data is burdened with considerable uncertainty. This study reports an original methodology that serves to demystify array results, reduce misinterpretation, and accurately quantify position and strength of acoustic sources. Traditional array results represent noise sources that are convolved with array beamform response functions, which depend on array geometry, size (with respect to source position and distributions), and frequency. The Deconvolution Approach for the Mapping of Acoustic Sources (DAMAS) method removes beamforming characteristics from output presentations. A unique linear system of equations accounts for reciprocal influence at different locations over the array survey region. It makes no assumption beyond the traditional processing assumption of statistically independent noise sources. The full rank equations are solved with a new robust iterative method. DAMAS is quantitatively validated using archival data from a variety of prior high-lift airframe component noise studies, including flap edge/cove, trailing edge, leading edge, slat, and calibration sources. Presentations are explicit and straightforward, as the noise radiated from a region of interest is determined by simply summing the mean-squared values over that region. DAMAS can fully replace existing array processing and presentations methodology in most applications. It appears to dramatically increase the value of arrays to the field of experimental acoustics. *Senior Research Scientist, Aeroacoustics Branch, Associate Fellow AIAA. †Senior Research Scientist, Aeroacoustics Branch, Senior Member AIAA. Copyright

Extension of DAMAS Phased Array Processing for Spatial Coherence Determination (DAMAS-C)

The present study reports a new development of the DAMAS microphone phased array processing methodology that allows the determination and separation of coherent and incoherent noise source distributions. In 2004, a Deconvolution Approach for the Mapping of Acoustic Sources (DAMAS) was developed which decoupled the array design and processing influence from the noise being measured, using a simple and robust algorithm. In 2005, three-dimensional applications of DAMAS were examined. DAMAS has been shown to render an unambiguous quantitative determination of acoustic source position and strength. However, an underlying premise of DAMAS, as well as that of classical array beamforming methodology, is that the noise regions under study are distributions of statistically independent sources. The present development, called DAMAS-C, extends the basic approach to include coherence definition between noise sources. The solutions incorporate cross-beamforming array measurements over the survey region. While the resulting inverse problem can be large and the iteration solution computationally demanding, it solves problems no other technique can approach. DAMAS-C is validated using noise source simulations and is applied to airframe flap noise test results.

Stochastic estimation of a separated-flow field using wall-pressure-array measurements

Concurrent, surface-pressure and planar, particle image velocimetry (PIV) measurements were obtained in the separating/reattaching flow region downstream of an axisymmetric, backward-facing step at a Reynolds number of 8081, based on step height. The surface-pressure and PIV measurements were used to investigate the evolution of coherent structures in the flow field by employing proper orthogonal decomposition (POD) and multipoint, linear, stochastic estimation (mLSE) analysis techniques. POD was used to determine the dominant modes in the pressure signature, while mLSE was used to estimate the dominant flow structures above the wall from the wall-pressure POD modes over a series of time steps. It was found that a large-scale, coherent structure develops in place (i.e., temporally) at approximately half the reattachment distance. Once this structure reaches a height equivalent to the step, it sheds and accelerates downstream. This growth in place, and then shedding, resembles the evolution of the flow str...

Aeroacoustic Measurements of a Wing-Flap Configuration

Aeroacoustic measurements are being conducted to investigate the mechanisms of sound generation in high-lift wing configurations, and initial results are presented. The model is approximately 6 percent of a full scale configuration, and consists of a main element NACA 63 \sub{2}-215 wing section and a 30 percent chord half-span flap. Flow speeds up to Mach 0.17 are tested at Reynolds number up to approximately 1.7 million. Results are presented for a main element at a 16 degree angle of attack, and flap deflection angles of 29 and 39 degrees. The measurement systems developed for this test include two directional arrays used to localize and characterize the noise sources, and an array of unsteady surface pressure transducers used to characterize wave number spectra and correlate with acoustic measurements. Sound source localization maps show that locally dominant noise sources exist on the flap-side edge. The spectral distribution of the noise sources along the flap-side edge shows a decrease in frequency of the locally dominant noise source with increasing distance downstream of the flap leading edge. Spectra are presented which show general spectral characteristics of Strouhal dependent flow-surface interaction noise. However, the appearance of multiple broadband tonal features at high frequency indicates the presence of aeroacoustic phenomenon following different scaling characteristics. The scaling of the high frequency aeroacoustic phenomenon is found to be different for the two flap deflection angles tested. Unsteady surface pressure measurements in the vicinity of the flap edge show high coherence levels between adjacent sensors on the flap-side edge and on the flap edge upper surface in a region which corresponds closely to where the flap-side edge vortex begins to spill over to the flap upper surface. The frequency ranges where these high levels of coherence occur on the flap surface are consistent with the frequency ranges in which dominant features appear in far field acoustic spectra. The consistency of strongly correlated unsteady surface pressures and far field pressure fluctuations suggests the importance of regions on the flap edge in generating sound.

Velocity and thermal structure, and strain-induced extinction of 14 to 100% hydrogen-air counterflow diffusion flames

Extensive results from axisymmetric convergent-nozzle and straight-tube opposed jet burners (OJBs) characterized strain-induced extinction of unanchored (free-floating), laminar H 2 /N 2 -air flames. Parameters included (a) plug-flow and parabolic input velocity profiles, (b) jet exit diameters ranging 2.7 to 7.2 mm for nozzles and 1.8 to 10 mm for tubes, (c) various relative jet gaps, and (d) 14 to 100% H 2 in the fuel jet. Extinction, a sudden rupture (blowoff) of the mostly-airside disk flame, occurred as fuel and air flows were slowly increased and a critical radial strain rate was exceeded. The disk flame was restored at much lower flows, unique to H 2 systems. Focusing schlieren, thermocouple, and airside LDV (and PIV) data confirmed the one-dimensional (I-D) character of nozzle-OJB flow fields; axial widths of velocity- and thermal-layers varied as (input strain rate) −1/2 for both nozzles and tubes. The global approximation of a I-D applied stress rate (ASR), using average air jet velocity divided by exit diameter, enabled high quality correlations of extinction data with varied H 2 concentrations, for both nozzles and tubes. Pre-extinction ASRs for nozzles agreed closely with LDV-measured centerline input strain rates; for tubes, however, an empirical factor of 3 produced close agreement. For methane-air extinction, nozzle-OJB ASRs agreed within 4% of independent nozzle and Tsuji burner results. For extinction of 100% H 2 -air, an ASR of 5670 1/s compared with 7350, 8140, and 8060 from independent 1-D numerical evaluations using potential-flow inputs; for 50 to 14% H 2 inputs, agreement was much closer. The nozzle-ASR/tube-ASR ratio for extinction was ≥3 for 2 inputs, 2.74 ± 0.03 for 50 to 100% H 2 inputs, and 2.83 for methane-air. Because these ratios exceeded 2.0, which “accounted” for centerline velocity inputs from parabolic profiles, an additional 3/2 radial strain component was apparent and was supported by the axial velocity gradient measurements.

Three-Dimensional Application of DAMAS Methodology for Aeroacoustic Noise Source Definition

At the 2004 AIAA/CEAS Aeroacoustic Conference, a breakthrough in acoustic microphone array technology was reported by the authors. A Deconvolution Approach for the Mapping of Acoustic Sources (DAMAS) was developed which decouples the array design and processing influence from the noise being measured, using a simple and robust algorithm. For several prior airframe noise studies, it was shown to permit an unambiguous and accurate determination of acoustic source position and strength. As a follow-on effort, this paper examines the technique for three-dimensional (3D) applications. First, the beamforming ability for arrays, of different size and design, to focus longitudinally and laterally is examined for a range of source positions and frequency. Advantage is found for larger array designs with higher density microphone distributions towards the center. After defining a 3D grid generalized with respect to the array’s beamforming characteristics, DAMAS is employed in simulated and experimental noise test cases. It is found that spatial resolution is much less sharp in the longitudinal direction in front of the array compared to side-to-side lateral resolution. 3D DAMAS becomes useful for sufficiently large arrays at sufficiently high frequency. But, such can be a challenge to computational capabilities, with regard to the required expanse and number of grid points. Also, larger arrays can strain basic physical modeling assumptions that DAMAS and all traditional array methodologies use. An important experimental result is that turbulent shear layers can negatively impact attainable beamforming resolution. Still, the usefulness of 3D DAMAS is demonstrated by the measurement of landing gear noise source distributions in a difficult hard-wall wind tunnel environment.

Noise spectra and directivity for a scale-model landing gear

An extensive experimental study has been conducted to acquire detailed noise spectra and directivity data for a high-fidelity, 6.3%-scale, Boeing 777 main landing gear. The measurements were conducted in the NASA Langley Quiet Flow Facility using a 41-microphone directional array system positioned at a range of polar and azimuthal observer angles with respect to the model. DAMAS (Deconvolution Approach for the Mapping of Acoustic Sources) array processing as well as straightforward individual microphone processing were employed to compile unique flyover and sideline directivity databases for a range of freestream Mach numbers (0.11 – 0.17) covering typical approach conditions. Comprehensive corrections were applied to the test data to account for shear layer ray path and amplitude variations. This allowed proper beamforming at different measurement orientations, as well as directivity presentation in free-field emission coordinates. Four different configurations of the landing gear were tested: a baseline configuration with and without an attached side door, and a noise reduction concept “toboggan” truck fairing with and without side door. Spectral analyses demonstrated that individual microphones could establish model spectra. This finding permitted the determination of unique, spatially-detailed directivity contours of spectral band levels over a hemispherical surface. Spectral scaling for the baseline model confirmed that the acoustic intensity scaled with the expected sixth-power of the Mach number. A comparison of spectra and directivity between the baseline gear and the gear with an attached toboggan indicated that the toboggan fairing may be of some value in reducing gear noise over particular frequency ranges. Finally, spectrum results were scaled quantitatively to data obtained from flyover tests of a full-scale 777 aircraft. The comparison showed good agreement in frequency and level, suggesting the general applicability of the model-scale noise results.

DAMAS Processing for a Phased Array Study in the NASA Langley Jet Noise Laboratory

A jet noise measurement study was conducted using a phased microphone array system for a range of jet nozzle configurations and flow conditions. The test effort included convergent and convergent/divergent single flow nozzles, as well as conventional and chevron dual-flow core and fan configurations. Cold jets were tested with and without wind tunnel co-flow, whereas, hot jets were tested only with co-flow. The intent of the measurement effort was to allow evaluation of new phased array technologies for their ability to separate and quantify distributions of jet noise sources. In the present paper, the array post-processing method focused upon is DAMAS (Deconvolution Approach for the Mapping of Acoustic Sources) for the quantitative determination of spatial distributions of noise sources. Jet noise is highly complex with stationary and convecting noise sources, convecting flows that are the sources themselves, and shock-related and screech noise for supersonic flow. The analysis presented in this paper addresses some processing details with DAMAS, for the array positioned at 90ϒ (normal) to the jet. The paper demonstrates the applicability of DAMAS and how it indicates when strong coherence is present. Also, a new approach to calibrating the array focus and position is introduced and demonstrated.

Aeroacoustic Evaluation of Flap and Landing Gear Noise Reduction Concepts

Aeroacoustic measurements for a semi-span, 18% scale, high-fidelity Gulfstream aircraft model are presented. The model was used as a test bed to conduct detailed studies of flap and main landing gear noise sources and to determine the effectiveness of numerous noise mitigation concepts. Using a traversing microphone array in the flyover direction, an extensive set of acoustic data was obtained in the NASA Langley Research Center 14by 22-Foot Subsonic Tunnel with the facility in the acoustically treated openwall (jet) mode. Most of the information was acquired with the model in a landing configuration with the flap deflected 39o and the main landing gear alternately installed and removed. Data were obtained at Mach numbers of 0.16, 0.20, and 0.24 over directivity angles between 56o and 116o, with 90o representing the overhead direction. Measured acoustic spectra showed that several of the tested flap noise reduction concepts decrease the sound pressure levels by 2 – 4 dB over the entire frequency range at all directivity angles. Slightly lower levels of noise reduction from the main landing gear were obtained through the simultaneous application of various gear devices. Measured aerodynamic forces indicated that the tested gear/flap noise abatement technologies have a negligible impact on the aerodynamic performance of the aircraft model.