Christopher S. Allen

EFFECT OF SURFACE TREATMENT ON ARRAY MICROPHONE SELF-NOISE

A method for reducing the flow-induced self-noise of acoustic array microphones was demonstrated in a wind tunnel. Unsteady boundary-layer flow in the wind tunnel induces large pressure fluctuations on exposed, flushmounted microphone diaphragms, reducing the signal-tonoise capability of microphone arrays. Two important steps were taken to reduce this background noise. First, the microphones were recessed to separate the microphones from the unsteady flow. Second, a porous surface material was placed above the microphones to act as an aerodynamic surface while allowing acoustic signals to pass through to the microphones. Previous attempts at this approach used perforated plates as the surface material, which tended to fatigue in the unsteady flow. The increased acoustic impedance of thicker materials caused reverberation between the surface and the microphone mounting plate. This latest attempt used a stretched sheet of Kevlar ® as the surface. The extreme strength and durability of the Kevlar ® withstood flow-induced fatigue while providing very low acoustic impedance with little attenuation of sound for most frequencies. Data is presented for two wind tunnel tests that demonstrate the capabilities of the recessed Kevlar ® array.

Flight Effects on the Far-Field Noise of a Heated Supersonic Jet

The influence of forward flight on the far-field noise of an underexpanded heated supersonic jet has been studied experimentally with a 12.5-cm-diam convergent nozzle operated in the NASA Ames Research Center 12.2 x 24.4 m (40 x 80 ft) wind tunnel. The nozzle was operated at nozzle pressure ratios up to 4.5 and stagnation temperature ratios from 2.45 to 3.45. The resulting velocity (based on fully expanded condition) range is from 586 to 858 m/s. The freestream Mach number was varied from 0 to 0.32. Far-field narrow band spectra were obtained at angles (measured from the inlet axis) covering a range from 30 to 155 deg. A small amplification of the overall sound-pressure level (2 dB) due to forward flight is observed in the forward quadrant. The mixing noise reduction in the aft quadrant due to forward flight is much smaller than that observed in corresponding cold jets.

Microphone Measurements In and Out of Airstream

The wind tunnel has become an important research facility for the study of aircraft and automobile noise. In this chapter, the acoustic characteristics of wind tunnels are discussed along with methods for conducting research in such an environment. Microphone measurements require low background noise and minimal reflections for accurate results. Typical sources of wind tunnel background noise are described including noise from the wind tunnel components, apparatus support struts, and microphones installed in the flow. In most cases, proper design of wind tunnel components and test apparatus are critical to successful aeroacoustic measurements. And, it is often necessary to add silencers and acoustic treatment to the facility. Criteria for proper simulation of aeroacoustic phenomena are discussed along with necessary data manipulations to correct for propagation effects, scaling to the correct source size, and extrapolating to the desired flight or drive-by situation. Finally, current methods are discussed for identification and analysis of noise sources using advanced signal analysis techniques.

Design of a deep acoustic lining for the 40- by 80-foot Wind Tunnel test section

‘Y Member AIAA * Aerospace Computing Inc., Senior tiember AlAA Fiberglass wedges were installed in the cavity czated between the flow boundary and pressure shell. The wedges are protected from the airflow by a porous interface at the flow boundary. The resulting acoustic lining is nominally 1.07 m deep except for certain shallow areas over structural beams, turntable apparatus, and.diffuser inlet. The deep lining encircles the 25.3-m long test section and is joined to a 15.2-cm deep shallow lining that extends 6.1 m into the diffuser inlet for a total treated length of 31.4 m streamwise. The lining is f&d to the original wind tunnel duct at the entry and exit of the test section.