Aeroacoustic Measurements

Acoustic wind tunnel test is an important way to study the noise mechanism of helicopter rotors. The broadband dipole noise distribution of helicopter rotors can be obtained from the acoustic wind tunnel measurement, which can be extremely helpful in analyzing the pulsation noise mechanism of rotor blade surfaces. However, during acoustic imaging of broadband dipole sources, there is often a problem of unwanted noise interference from other experimental setups of the wind tunnel, such as the rotor shaft. To accurately visualize the broadband dipole sound sources, in this paper, a moving dipole-type sources beamforming method suitable for the wind tunnel environment is proposed. The reduced-rank cyclic Wiener filter is firstly developed to extract the rotor broadband noise signal in wind tunnel tests; a beamforming method for moving dipole sound sources is then proposed according to the characteristics of the generation and propagation of the dipole sound sources. The BO105 helicopter rotor model is tested in the FL-17 aeroacoustic wind tunnel of the China Aerodynamics Research and Development Center (CADRC). The wind tunnel test results demonstrate that the method can effectively suppress noise interference in the wind tunnel and accurately extract and visualize broadband dipole sources of rotor blade surfaces.

Wind tunnel measurement quality in testing of a standard model

Chapter 12 - Phased arrays

A low-speed closed-loop wind tunnel with an open-jet test section having a maximum windspeed of 60 m/s is being modified for future aero-acoustic testing. The wind tunnel is new to Carleton University as of 2021 and no acoustic characterization tests have been completed until recently. The objective of this research is to acoustically characterize the wind tunnel before and stepwise through its acoustic modifications. Multiple tests have been made using ¼-inch microphones at different locations inside the plenum to record the existing out-of-flow background noise. The first planned wind tunnel modification is a height extension to the current plenum with dimensions 0.74 × 1.8 × 1.3 m (H × L × W) to increase the observer distance for the microphones. Once the plenum volume is increased, acoustic treatment will be applied to the interior of the plenum to create a hemi-anechoic measurement environment. An in-flow microphone mount was designed and manufactured with computer-controlled vertical traversing; it will be installed after the plenum modifications are complete.

Model scale tests of modern ‘open rotor’ propulsor concepts that have potential for significant fuel burn reduction for aircraft applications were completed at NASA Glenn Research Center. The recent test campaign was a collaboration between NASA, FAA, and General Electric (GE). GE was the primary industrial partner, but other organisations were involved such as Boeing and Airbus who provided additional hardware for fuselage simulations. The open rotor is a modern version of the UnDucted Fan (UDF®) that was flight tested in the late 1980s through a partnership between NASA and GE. The UDF® was memorable for its scimitar shaped propeller blades and its unique noise signature. Design methods of the time were not able to optimise for both high aerodynamic efficiency and low noise simultaneously. Contemporary CFD/CAA based design methods can produce open rotor blade designs that maintain efficiency with acceptable acoustic signatures. Tests of two generations of new open rotor designs were conducted in the 9’ × 15’ Low Speed Wind Tunnel and the 8’ × 6’ Supersonic Wind Tunnel starting in late 2009 and completed in early 2012. Aerodynamic performance and acoustic data were obtained for take-off, approach and cruise conditions in isolated and semi-installed configurations. Additional detailed flow diagnostic measurements and acoustic measurements, including canonical shielding configurations, were obtained by NASA. NASA and GE conducted joint systems analysis to evaluate the performance of the new blade designs on a Boeing 737 class aircraft. The program demonstrated a 2-3% improvement in overall net efficiency relative to the best efficiency designs of the 1980s while nominally achieving 15-17 EPNdB noise margin to Chapter 4 (at a Technology Readiness Level of 5) for a notional aircraft system defined by NASA.

The present study is an attempt to computationally analyze major and minor energy losses in a low speed wind tunnel (LSWT) with special emphasis on pressure drop characterization together with a co-relating theoretical analysis. This study produces preliminary results for analysis and design of an open-circuit LSWT which is to be constructed at Center of Fluid Mechanics at Mondragon University (MU) for analyzing low speed wind flow application(s) with reduced form pressures losses for flow past bluff bodies. Various losses at different sections in the wind tunnel are calculated and total pressure drop across the wind tunnel are estimated using computational fluid dynamics (CFD) and theoretical methods.

The Low-Speed Anechoic Wind Tunnel at Sun Yat-Sen University is a premier facility designed for experimental research on aerodynamics and aeroacoustics of aircraft scaled models or components. As a 3/4 closed-circuit wind tunnel with a contraction section nozzle dimension of 0.8 m × 0.6 m, it can be configured either as a closed test section or an open-jet section. The freestream velocity in the closed test section can be adjusted from 5 m/s to 80 m/s, while the openjet section ranges from 10 m/s to 70 m/s. The test section is enclosed within an anechoic chamber with dimensions of 5.5 m × 6 m × 3.5 m to ensure non-reflecting conditions. After the wind tunnel’s construction, comprehensive tests were conducted to evaluate its acoustical and aerodynamic performance. This paper presents the design criteria and characterization of the wind tunnel’s key components, along with the methods and results of benchmarking the flow quality and acoustic performance. This wind tunnel features exceptional flow quality with low turbulence intensity of 0.08% , 1% in the closed test section and open-jet section, respectively. The minimal level of background noise can attain a decibel level of 6.9 dBA under no-flow conditions, and at maximum flow velocity (70 m/s ) in the open-jet test section, the background noise reached 74.9 dB(A).The findings demonstrate that the aerodynamic and aeroacoustic performance indicators of the wind tunnel are comparable to similar facilities worldwide, positioning it as the leading aeroacoustic research facility in the Guangdong-Hong Kong-Macao Greater Bay Area at present.

Extensive outdoor and laboratory-scale experiments on sonic boom propagation in turbulent atmosphere have shown that shock wave amplitude and rise time are important parameters responsible for sonic boom annoyance. However, accurate measurement of the shock front structure with standard microphone remains a challenge due to the broadband spectrum of the N-wave shock front. In this work the experimental setup utilizing a spark source has been designed and built to investigate nonlinear N-wave propagation in homogeneous medium. Short duration (30μs) and high amplitude (1 kPa) spherically divergent N-waves were generated. In addition to acoustic measurements with 1/8" B&K microphones, the shadowgraphy method using short duration flash lamp (20 ns) and CCD camera was employed to assess the shock front structure at different distances from the spark. It is shown that the shock rise time measured by the shadowgraphy method was in a good agreement with the theoretical predictions and it was 10 times shorter than in microphone measurements. The widening of the shock in acoustic measurements was therefore due to the limited bandwidth of the microphone. The combination of modeling, acoustic and optical measurements provided an accurate calibration of the shock wave measuring system. [Work supported by RFBR and INTAS.]

This paper presents the design and construction of a small-scale low-speed open circuit wind tunnel suitable for experimental testing and analysis of aerodynamic phenomena. The wind tunnel is specifically designed for low-speed applications, focusing on flows within velocities of 35 meters per second with turbulence conditions below 10 percent. The design process of a wind tunnel involves considerations such as dimensions, design parameters, and instrumentation to ensure accurate flow measurements, with key components including a settling chamber, contraction section, test section, diffuser, and drive system working together to remove disturbances, create uniform flow, minimize blockage effects, and generate required airflow for accurate aerodynamic testing. This paper highlights the design considerations, key components, construction techniques, and the significance of the wind tunnel for aerodynamic testing, validation of simulations, and education. The designed wind tunnel is capable of producing controlled and repeatable low-speed flow conditions, making it suitable for a wide range of applications, including aerodynamic testing of scaled- down models, validation of Computational Fluid Dynamics (CFD) simulations, and educational purposes. The small-scale design allows for cost-effective construction and operation while maintaining reliable results.

In order to avoid the defects of contact measurement, such as limited range, complex constructing and disability of 3-D parameter acquisition, we built a binocular videogrammetric system for measuring 3-D geometry parameters of wind tunnel test models, for instance, displacement, rotation angle and vibration, in low-speed wind tunnel. The system is based on the principles of close-range digital photogrammetry. As a non-contact system, it acquires parameters without interference in the experiments, and it has adjustable range and simple structure. It is worth mentioning that this is a Realtime measurement system, so that it can greatly compress the experiment period, furthermore, it is also able to provide some specific experiments with parameters for online adjustment. In this system, images are acquired through two industrial digital cameras and a PCI-E image acquisition card, and they are processed in a PC. The two cameras are triggered by signals come from a function signal generator, so that images of different cameras will have good temporal synchronization to ensure the accuracy of 3-D reconstruction. A two-step stereo calibration technique using planar pattern developed by Zhengyou Zhang is used to calibrate these cameras. Results of wind tunnel test indicate that the system can provide displacement accuracy better than 0.1% and rotation angle accuracy better than 0.1 degree, besides, the vibration frequency accuracy is superior to 0.1Hz in the low-frequency range.

Testing otter board hydrodynamic performances in wind tunnel facilities

The historical development of laser velocimetry and its application to low-speed (less than 100 m/sec) aerodynamic flows in the subsonic wind tunnels at Ames Research Center is reviewed. A fully three dimensional velocimeter for the Ames 7- by 10-Foot Wind Tunnel is described, and its capabilities are presented through sample data from a recent experiment. Finally, a long-range (2.6 to 10 m) velocimeter that is designed to be installed within the test section of the Ames 40- by 80-Foot Wind Tunnel is described and sample data are presented.

The generation of a thick fully turbulent boundary layer is investigated in a lowspeed wind tunnel at a nominal zero pressure gradient over the Reynolds number range 0.145× 10 6 · Rex · 0.58× 10 6 . The wind tunnel floor natural boundary layer is laminar with thickness ± between 5.76 mm and 8.13 mm. Different tripping devices are tested to trigger transition so to double the boundary layer thickness and provide a fully established turbulent velocity profile. Using a trip wire significantly increases ± but leads to an unsatisfactory velocity profile. Using a sandpaper strip slightly increases ± but keeps the boundary layer laminar. Using a strip of sharp-edged silicon granules doubles the boundary layer thickness that increases up to 20 mm and the mean velocity profiles are a good fit to the logarithmic law of the wall over the outer region of the boundary layer. The spectral decay of turbulent kinetic energy in this outer layer is exponential and close to i5/3, indicating turbulence equilibrium. This work is of practical interest to wind tunnel practitioners for generating equilibrium thick turbulent boundary layers at low Reynolds numbers.

In this paper, low-speed smoke wind tunnel has been designed and fabricated for the insects’ flow field visualization. The test section and the contraction section of the tunnel are optimized and determined as to size by the method of computational fluid dynamics. And fairing devices are equipped in different sections to reduce the turbulence intensity and increase the flow uniformity in the experimental sections. For the smoke visualization of small insects, the smokeemitting equipment has been specially designed and carefully debugged. Composed of wind tunnel, light source and high-speed camera, experimental platform for visualization and filming of insect flight flow field has been established. B esides, the feasible and stable method for insect fixing has been designed. With the smoke wind tunnel, flow filed visualization experiment for the honeybee’s flapping was conducted and smoke flow filed in the experiment was recorded and analyzed. Near-filed and far-filed vortex structure when the honeybee fly can be recorded clearly. The expe rimental results indicate that the experimental platform is appropriate for flow filed study on insects flapping. Keyw ords: wind tunnel; honeybee; smoke lines; flow field

Reduction of 3/4 open jet low-frequency fluctuations in the S2A wind tunnel