Validation of a Low Reynolds Number Aerodynamic Characterization Facility

At present, there is a great interest in the development of new airfoils for wind turbines and high-lift wings of unmanned aerial vehicles (UAV). The requirements for such airfoils differ from conventional aircraft airfoils, because of structural reasons and extreme operating conditions. So, wind turbine airfoils operate frequently under fully separated flow when stall is used for power regulation at high wind speeds. At the same time design of airfoils for wings UAV poses the problem of availability of high-lift at low Reynolds number. Modern airfoils are to a large extent developed from numerical methods. However, the complex flow conditions such as separation at high angles of attack, laminar separation bubbles and the transition from laminar to turbulent flow are difficult to predict accurately. Hence, testing of airfoils at a two-dimensional condition is an important phase in airfoil design. The development and validation of a 2D testing facility for investigation of single and multi-element airfoils in the wind tunnel Т-102 with open test section are considered in this article. T-102 is a continuous-operation, closed-layout wind tunnel with two reverse channels. The test section has an elliptical cross-section of 4 ×2,33 m and a length of 4 m. Two big flat panels of the L × H=3 ×3,9 m size installed upright on balance frame aligned with the free stream are used for simulating two-dimensional flow in the tunnel test section. The airfoil section in the layout of a rectangular wing is mounted horizontally between flat panels with minimum gaps to ensure 2D flow conditions. The aerodynamic forces and pitch moment acting on the model were measured by wind tunnel balance. To determine boundary corrections for a new test section of wind tunnel, the experimental investigation of three geometrically similar models has been executed. The use of boundary corrections has provided good correlation of the test data of airfoil NACA 6712 with the results obtained from the wind tunnel except for lift and drag coefficient values at high angles of attack.

A detailed flow survey of the test section of the Langley Full-Scale Tunnel is described. Flow angularity, dynamic pressure, and turbulence intensity were measured in the test section volume over a matrix with three foot spacing. Instrumentation consisted of a rake of four five hole probes and two hot-wire anemometers. Details of probe calibration, installation, and use are presented in addition to survey results. Facility Description Old Dominion University (ODU), working under a Memorandum of Agreement with NASA Langley Research Center, has operated the Langley Full-Scale Tunnel (LFST) since 1997. This facility is the second largest in the United States in terms of test section size and is the largest university run wind tunnel in the world. During its seventy years of service, a wide variety of aerospace test programs have been undertaken. Sub-scale model testing has included high angle of attack testing of current U.S. front-line fighters, high-lift systems on supersonic and subsonic transports, parachutes, and even an airship, submarine, and building complex. Recently testing has included fullscale automobiles, trucks, and UAVs. The open jet test section of the LFST (formerly the NASA 30 by 60) is semi-elliptical in cross section with a width of 60 ft and a height of 30 ft as shown in figure I. The elevated ground board is 42.5 ft wide by 52.3 ft long and features a turntable with a diameter of 28.5 ft. * Graduate Research Assistant Student Member AIAA t Associate Professor of Aerospace Engineering, Senior Member AIAA Copyright © 2002 by the American Institute of Aeronautics and Astronautics, Inc. Power is supplied by two 4000 HP electric motors driving two 35.5 ft diameter fourbladed fans. The dual return tunnel circuit design has a nozzle contraction ratio of approximately 5:1, and no flow conditioning devices such as honeycombs or screens anywhere in the circuit. The test section ground board and balance shield represent the only major modifications to the circuit since the commissioning of the facility. Introduction The large test section and open jet design are ideal for testing large models and UAVs with negligible boundary effects. Special vents provide fresh air exchange, useful when testing with power on. There are two model aircraft support systems available depending on size and weight. Both systems allow a 100° angle of attack sweep and a ±90° yaw sweep. A large scale x-y-z traverse (survey carriage) can be used to position a second model or survey probes in the test section. Interest in a modern flow survey originally stemmed from a planned NASA Blended Wing Body (BWB) test using a large scale model. There have been several flow surveys in the history of the tunnel and much folklore surrounding the flow quality on the tunnel centerline.'' The merging streams of the dual return design were thought to have created a measureable shear layer on the centerline with an elevated turbulence intensity. This report describes the flow quality in the Langley Full-Scale Tunnel (LFST) in terms of flow angularity, dynamic pressure and turbulence intensity. Aerodynamic flow field surveys were made over the entire test section volume. The instrumentation consisted of four five hole hemispherical-head probes for dynamic pressure and flow angularity measurements, two constant-temperature hot-wire anemometers for 1 American Institute of Aeronautics and Astronautics (c)2002 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization.

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.

The present study aims to investigate the effect of the contraction profile on the flow quality such as the turbulent intensity, flow uniformity and pressure compared to baseline case (existing one). Primarily, losses of every wind tunnel part should be calculated. Hence, the total pressure losses of the whole wind tunnel circuit can be determined and all these losses are added up. This technique is required to evaluate the required power to operate the wind tunnel. As result, this method is well-suited with both of opened and closed-circuit wind tunnels. Moreover, two different profiles of the contraction are compared to the baseline wind tunnel (existing one) using the Computational Fluid Dynamics (CFD) as a basis for the comparison and analysis of the airflow properties through the whole wind tunnel design. The necessary parameters are the flow conditions within the test section that included the velocity, turbulent intensity and flow uniformity as well as the pressure distribution. The results showed that the turbulent intensity is decreased by 29.7 % with more uniform flow inside the optimized wind tunnel, when the “eye design contraction wall” is used. The contraction profile (wall design shape) plays an important role to effect on the flow quality inside the test section. Therefore, a special attention should be given when designing the wind tunnel with a specific contraction profile.

A novel type of semi-active jet turbulence grid

In linear cascade wind tunnel tests, a high level of pitchwise periodicity is desirable to reproduce the azimuthal periodicity in the stage of an axial compressor or turbine. Transonic tests in a cascade wind tunnel with open jet boundaries have been shown to suffer from spurious waves, reflected at the jet boundary, that compromise the flow periodicity in pitch. This problem can be tackled by placing at this boundary a slotted tailboard with a specific wall void ratio s and pitch angle α. The optimal value of the s-α pair depends on the test section geometry and on the tunnel running conditions. An inviscid two-dimensional numerical method has been developed to predict transonic linear cascade flows, with and without a tailboard, and quantify the nonperiodicity in the discharge. This method includes a new computational boundary condition to model the effects of the tailboard slots on the cascade interior flow. This method has been applied to a six-blade turbine nozzle cascade, transonically tested at the University of Leicester. The numerical results identified a specific slotted tailboard geometry, able to minimize the spurious reflected waves and regain some pitchwise flow periodicity. The wind tunnel open jet test section was redesigned accordingly. Pressure measurements at the cascade outlet and synchronous spark schlieren visualization of the test section, with and without the optimized slotted tailboard, have confirmed the gain in pitchwise periodicity predicted by the numerical model.

For a number of manmade and natural flyers the aerodynamics of low Reynolds number plays a crucial role. Active study of birds and insects is growing rapidly for development of Micro Air Vehicles (MAVs). MAVs are capable of performing tasks like surveillance, environmental monitoring and threat assessment in hostile environment. When compared to general civilian aircrafts MAVs fly at low Reynolds number regime of 105 or lower. A considerable change of aerodynamic characteristics is seen between aircrafts that operate in low and high Reynolds number regimes. Low Reynolds number flyers are sensitive to wing gust and deformation because of being lightweight and operating at low speeds. The birds have nature’s finest locomotion system when it comes to maneuvering effectively through space. This review paper will assist further studies in the field of low Reynolds number flyers that can be used for future development on MAVs. A holistic review of past two decades of research work and studies done on low Reynolds number flyers has been presented in this paper. The Theoretical, Experimental and Numerical techniques followed by researchers have been highlighted so that the study on biological and manmade flyers can be systematically presented. It was found that researchers preferred Numerical and Experimental techniques while conducting aerodynamic analysis of low Reynolds number flyers. An attempt has been made in this paper to fill the research gap that is there in mimicking the flight of animals and thus developing man-made systems having higher efficiency.

The present study intends to investigate the potentiality of the single dielectric barrier discharge plasma actuators (SDBDPAs) to reattach the separated flow, occurring at low Reynolds numbers. For this aim, a curved wall plate, which profile shape was designed to reproduce the suction surface of a low-pressure turbine (LPT) blade, was installed in the test section of a closed loop wind tunnel and a SDBDPA was placed in a groove made over it, at the front of the adverse pressure gradient region. The flow behavior in absence of actuation has been experimentally investigated by particle image velocimetry (PIV) and laser Doppler velocimetry measurements (LDV). Moreover, sinusoidal voltage excitation with amplitude of 8 kV and frequency of 2 kHz was applied to the SDBDPA and PIV measurements were also performed in presence of actuation. The applied voltage and the discharge current have also been recorded simultaneously, and they have been used for the determination of the device dissipated power. Different wind tunnel free stream velocities were investigated, both in absence and presence of actuation. The effect of the active flow control was then studied in the entire measurement domain by analyzing the velocity fields, the turbulence intensity (Tu) values, the momentum coefficient (cμ) and the boundary layer shape factor (H12). In absence of actuation, a large reverse flow and a turbulence intensity up to ≈60% was observed in the separation region. Good agreement was found between the flow results obtained by the two velocity measurement techniques. Considering the actuated cases, it was found that, in all the tested operating conditions, actuation implied a reduction of the separated region and the turbulence intensity, even if a low flow control effect was noticed at the highest tested velocity.

<div class="htmlview paragraph">The paper presents new data to document the effective test area in a model ¾ open jet wind tunnel with a long test section using drag measurements. The drag on MIRA Pickup models (up to 20% blockage) was tested and found to be virtually independent of blockage when tested in a non-resonant test section with L/D ratio of about 8. An analysis of these test results is presented to predict the required test section L/D as a function of blockage. In addition, the effects of plenum dimensions are presented. Based on these results, the test sections of present open jet test sections have a limited range for testing models without corrections for the effects of blockage, model location, and test section length. A longer test section enables direct measurement of drag even with high blockage models. Testing with higher blockage can mean lower test facility costs.</div> <div class="htmlview paragraph">The test section design technology described in this report is the subject of a pending patent application.</div>

Until recently, the problem of low test Reynolds numbers has limited the usefulness of wind tunnels, especially at transonic speeds. Perhaps the best solution to the problem of low Reynolds number comes by operating a relatively big tunnel (2 × 2 m test section or larger) at a relatively high pressure (up to 5 bar) and at cryogenic temperatures using nitrogen (or nitrogen-rich air) as the test gas. The first cryogenic wind tunnel was a low-speed tunnel built at the NASA Langley Research Center in 1972. Since then, workers at research centers around the world have started a large number of cryogenic wind-tunnel projects. The degree of success of the various projects depends on many things. The more successful cryogenic tunnels reflect a skillful combination of cryogenic technology with wind-tunnel technology. I briefly review the principles of operation and the history of cryogenic wind tunnels. I describe some of the more significant cryogenic wind-tunnel projects not covered by other papers at this conference. I also describe some of the problems we must solve before we can realize the full potential of cryogenic wind tunnels. Finally, I briefly describe the heavy-gas option for high Reynolds number testing.

Effect of turbulent inflows on airfoil performance for a Horizontal Axis Wind Turbine at low Reynolds numbers (part I: Static pressure measurement)

Particle Image Velocimetry (PIV) relies upon the introduction of particle tracers that scatter sufficient light and follow the flow accurately. The use of submillimetre helium-filled soap bubbles (HFSB) as flow tracers for PIV is investigated for the purpose of enabling velocity measurements in large-scale industrial wind tunnels. That soap bubbles reflect more light than scattered by small liquid droplets or solid particles, allowing larger volumes to be illuminated, is a long known fact and has caught the attention of aerodynamicists since the 1930s. The difficulty encountered during initial efforts on using soap bubbles for accurate measurements revolves around the lack of control during the generation of these tracers, and the failure in presenting evidence that they could accurately follow the flow. Proof of concept that HFSB could be used for accurate flow measurements in wind tunnels was presented in the year that preceded the beginning of this work. In this thesis, the generation and control of HFSB and their tracing fidelity are studied through a series of experiments and simulations, bringing large-scale PIV using HFSB to the technology maturity level required for industrial measurements. High-speed shadowgraphy at the bubble generator exit revealed the main regimes of bubble generation. A regular, periodic and controlled generation bubbles of monodisperse size distribution, namely, the bubbling regime, was obtained by properly tuning of the input flow rates. The relation of the later with the bubble size and production rate was also obtained from these visualizations. Measurements of the HFSB velocity in the stagnation region ahead of a cylinder, obtained with Particle Tracking Velocimetry (PTV), relative to the flow velocity (slip velocity) were used to retrieve the HFSB time response and the ratio of helium to soap flow rates that satisfy the neutral buoyancy condition, in which the soap bubble density equals that of the surrounding air flow. Simulations of the particle motion in a rectilinear oscillatory flow was used to quantify the importance of the unsteady forces acting on a particle and to derive empirical relations for estimating the HFSB slip velocity in flows where the unsteady forces are relevant. In this case, the particle slip velocity is shown to depend on three parameters: the particle Reynolds number, the ratio of particle-to-fluid density and the flow time-scale. These cannot be combined into a single non-dimensional Stokes number. The validity of the empirical relations were extended for the analysis of the slip velocity of a particle travelling around an object. Based on the later, a method for deriving the density of a nearly-neutrally-buoyant particle that comprises the effects of unsteady forces and allows mismatch of acceleration between the particle and the flow was described. The tools developed for slip velocity analysis using the simulations were applied to assess experimental data from large-scale PIV measurements performed at the Low-Speed Tunnel (LST) of the German-Dutch Wind Tunnels (DNW). The experiments were realized in the flow around an airfoil of 70 cm chord at free stream velocity up to 70 m/s, reaching a chord-based Reynolds number of 3.2 million. PIV measurements using HFSB at this speed and Reynolds number were unprecedented. The results have indicated variations of the bubble density (20-30%) occurring post-generation. The tracing fidelity of HFSB in wall-bounded turbulence is investigated by comparing measurements in a turbulent-boundary layer of the mean velocity and Reynolds stress profiles, with those obtained with micrometre oil droplets (reference) and submillimetre air-filled soap bubbles (AFSB). The results have shown that the statistics of the first and second moments of velocity are well captured by all three investigated tracers, even by the heavier-than-air AFSB, which were shown to be poor tracers in the stagnation of a cylinder. Mechanisms of preferential concentration in turbulence were attributed as the cause of the better traceability observed. The thesis is concluded with a successful industrial application in the Large Low-Speed Facility (LLF) of DNW (9.5×9.5 m2 test section) around a tiltrotor aircraft in three flight modes, hover, transition and cruise, and tunnel speeds up to 60 m/s. The bubbles were introduced into the flow using a 3×3 m2 seeding rake, containing 400 bubble generators. The PIV measurements were performed in stereoscopic configuration in a field-of-view of 1.1×1.1 m2.

The effects of unsteady wind on the performances of a newly developed cross-axis wind turbine: A wind tunnel study

<div class="htmlview paragraph">The boundary interference of high blockage models inside open jet test section is studied in three phases. First, a wind tunnel test was performed using a high blockage automotive model inside the 1/15<sup>th</sup> scale Langley Full Scale Wind tunnel. Second, a CFD simulation was done using CFL3D code (developed by NASA Langley Research Center). Finally, a panel method was used to assess the boundary interference and to study the effect of the collector. The objective of the study is to highlight the challenges in assessing the boundary interference for high blockage models. A secondary object is to present a model to integrate all the available information from the wind tunnel test and the CFD simulation to solve the problem using a panel method.</div>

This study investigated the effects of using bionics in reducing the low-frequency fluctuation phenomenon in three-fourths open-jet automobile wind tunnels. Low-frequency fluctuation in these wind tunnels can cause pulsations of velocity and pressure in the test section, leading to measurement accuracy reduction. The unique shape of the seal vibrissa had been demonstrated to effectively suppress vortex-induced vibrations. Here the wind tunnel nozzle was optimized with seal-vibrissa-shaped protrusions. The effects and corresponding mechanism of seal-vibrissa-shaped nozzle toward jet shear layer vortex structures and low-frequency fluctuation phenomenon were analyzed. A 1:15 scaled three-fourths open-jet wind tunnel was used in the study. The results indicated that after the implementation of the seal-vibrissa-shaped nozzle, the turbulence intensity was decreased at the bottom of the test section and the pressure and velocity fluctuations were effectively suppressed. The three-dimensional vortex structures separating from the surface of vibrissa destroyed the original two-dimensional vortex structures produced by the original nozzle. Many small vortex structures with dispersed frequencies and energies were formed, which decreased the low-frequency fluctuations. The study showed that the seal-vibrissa-shaped nozzle had the potential of inhibiting the large-scale vortex separation in the jet shear layer, which could effectively reduce the low-frequency fluctuations.