The recent interest in the development of small unmanned aerial vehicles (UAVs) and micro air vehicles has revealed a need for a more thorough understanding of the aerodynamics of small airplanes flying at low speeds. In response to this need, a study of the lift, drag, and pitching moment characteristics of wings of low aspect ratio operating at low Reynolds numbers are presented. Wind-tunnel tests of wings with aspect ratios between 0.5 and 2.0, four distinct planforms, thickness-to-chord ratios of ≈ 2%, and 5-to-1 elliptical leading edges have been conducted as part of this research. The Reynolds numbers considered were in the range of 7 × × 10 4 to 2 × × 10 5 . Analysis of the data includes comparison of lift-curve slope, nonlinear equation approximations, maximum lift coefficient, and center of lift.

9R7. Aeroacoustic Measurements. - Edited by TJ Mueller (Dept of Aero and Mech Eng, Univ of Notre Dame, 112 Hessert Center, Notre Dame IN 46556-5684). Springer-Verlag, Berlin. 2002. 313 pp. ISBN 3-540-41757-5. $99.00.Reviewed by MG Prasad (Dept of Mech Eng, Stevens Inst of Tech, Hoboken NJ 07030).It is well known that experimental work plays an important role in acoustics. In particular, the acoustical measurements in the presence of airflow is challenging due to test conditions. This book deals with this challenging and important topic in acoustics namely aeroacoustics. The five chapters of the book deal with various aspects aeroacoustic measurements with applications. Chapter 1, on “Microphone measurements in and out of air stream,” deals with the acoustic characteristics of wind tunnels. The discussions include criteria for simulation of aeroacoustic test environment with requirements such as propagation effect, scaling, low background noise levels, etc. Chapter 2 presents “Beam forming in acoustic testing.” The use of beam forming technique and its applications in the modeling of wind tunnel noise with non-acoustic hard walls and free jets are discussed. Several beam forming algorithms with techniques for removing interference is discussed in this chapter. Chapter 3 covers “Aeroacoustic phased array testing in low speed wind tunnels.” A comprehensive treatment of the design and application of acoustic phased array measurement capability for low speed wind tunnels is presented. Various aspects of operation such as calibration, data handling, etc, are discussed. Chapter 4 is on “Source characterization by correlation techniques.” The characterization of acoustical sources using correlation techniques is presented. The chapter discusses various mathematical aspects through cross-correlation and coherence functions. Various examples including correlation measurements between surface pressure and far field acoustic pressure field data are included. The last chapter is “An anechoic facility for basic aeroacoustic research.” The chapter presents the design, construction, and performance evaluation of an anechoic wind tunnel facility for low speed and low turbulence applications. Use of such a facility to experimentally investigate fluid-solid interaction is discussed. A specific wind tunnel facility is described. As an example the use of the facility for a marine propeller response to inflow disturbance is discussed. All chapters are well presented. In summary, Aeroacoustic Measurements is a nice integration of presentations of various recognized contributors. The book deals with important issues in design and performance evaluation of wind tunnel test facilities, measurement techniques, and mathematical aspects. Nice photographs and figures are included. Each chapter has a large number of references. The book is recommended for libraries and all graduate students, researchers, and professionals in acoustics. In particular, those involved in experimental aeroacoustics will find this book very useful.

Results are presented of an experimental investigation into the aeroacoustic response characteristics of rotor turbulence ingestion. To fully characterize the rotor response, both the ingested velocity field and the resulting far-field sound were measured. The results are presented of a detailed velocity characterization, which was performed upstream of the rotor. The velocity measurements included an evaluation of the streamwise development of turbulence characteristics downstream of the grid, a high-resolution mapping of the spatial distribution of the mean velocity and rms turbulence fluctuations in the rotor inlet plane, and the development of a semi-empirical, functional representation of the three-dimensional wave number spectral density of the ingested turbulence. These data comprise a comprehensive empirical velocity model that was developed for specific application to rotor turbulence ingestion noise.

This is the second of two papers that discuss an experimental investigation of the aeroacoustic response characteristics of a 10-bladed rotor to grid-generated turbulence. To characterize empirically the rotor response, both the ingested velocity field and resulting far-field sound were measured. In part 1, an empirical velocity characterization of the grid- generated turbulence field was presented. This characterization culminated in a semi-empirical model of the ingested small-scale turbulence field and a modal decomposition of the circumferentially varying mean velocity field in the rotor inlet plane. This detailed velocity model is now used to investigate the aeroacoustic response of a 10-bladed rotor ingesting the grid-generated turbulence field. In particular, the semi-empirical turbulence model is used, in conjunction with theoretical spectral analysis techniques, to predict the far-field sound generated by the 10-bladed rotor. These predictions are compared to corresponding measured data to assess the fidelity of the spectral analysis methods and the semi-empirical turbulence model. Finally, the measured 10-bladed acoustic response is compared to the corresponding response of a 4-bladed rotor ingesting the same grid-generated turbulence field. These comparisons demonstrate the effect of geometry on the rotor acroacoustic sensitivity to both large-scale, spatial mean velocity modes and small-scale turbulence.

The presence of endplates (or sideplates) in two-dimensional wind-tunnel force measurements on airfoils has a strong effect on lift and drag coefficients at low Reynolds numbers. Results on an Eppler 61 airfoil indicate that the endplates are responsible for a sharp decrease in the airfoil performance. The lift coefficient is reduced and the drag coefficient is increased due to the interaction of the airfoil boundary layer with the sideplate boundary layer.

The effects of the vortical wake shed by a finite span canard on a low Reynolds number airfoil were examined. Aerodynamic performance was evaluated through direct measurements of lift, drag, and 1/4-chord pitching moment. Spanwise static pressure and surface film visualization data were also acquired. A reduction in the down-stream airfoil drag coefficient and an increase in its lift/drag were noted in the presence of the canard for a wide range of configurations

The formation and growth of transitional separation bubbles can significantly affect boundary-layer development on airfoils operating at low chord Reynolds numbers. Of primary concern is the change in boundary-layer thickness between laminar separation and turbulent reattachment. This can be estimated using semiempirical methods, such as the one devised by Horton (1968), which are based on solutions to the integral forms of the boundary-layer equations. The applicability of these methods at low Reynolds numbers was investigated using hot-wire measurements of bubbles formed on an NACA 66(3)-018 airfoil at chord Reynolds numbers of 50,000-200,000. The momentum thickness growth between separation and transition was found to be similar to that predicted for a laminar half-jet and appears to be influenced by the momentum thickness Reynolds number at separation. This parameter also was found to have a noticeable effect on the Reynolds number based on the length of a bubble's laminar portion.

Etude du comportement des bulles de decollement. Determination des distributions de pression et visualisation de l'ecoulement

An experimental study was conducted on a two-dimensional Wortmann FX63-137 airfoil model with a 13%-chord leading-edge flap and a 25%-chord trailing-edge flap at chord Reynolds numbers of 1.0 and 1.5 x 10. Deflecting the leading edge was shown to increase the maximum lift coefficient up to 9.2% while also increasing the maximum l/d. Q/Q experienced negligible improvement with leading-edge deflection. A leading-edge droop angle of about 5 deg was found to provide the best performance. Deflecting the trailing edge alone was shown to increase the maximum lift coefficient by up to 14.7%. Deflecting both the leading and trailing edges increased the maximum lift by up to 22.7%. Flow visualization revealed that drooping the leading edge moved the laminar separation bubble slightly downstream. Deflecting the trailing edge downward was shown to shift the line of turbulent separation toward the leading edge.

An experimental investigation was conducted in order to document the structure and behavior of laminar separation bubbles at low Reynolds numbers. Data of this type are necessary if the currently insufficient analytical and numerical models are to be improved. The laminar separation bubble that forms on a NACA 663-018 airfoil model was surveyed at chord Reynolds numbers of 50,000-200,000 at angles of attack of 8-12 deg. The effects of the various testing conditions on the separation bubble were isolated and the data were analyzed in relation to existing separation bubble correlations in order to test their low Reynolds number applicability. This analysis indicated that the chord Reynolds number and the disturbance environment strongly influence the experimental pressure distributions. These effects must be included in any analytic prediction technique applied to the low Reynolds number flight regime.

An experimental investigation was conducted on the performance and boundary layer characteristics of the Wortmann FX 63-137 airfoil with and without trip wire roughness. Data were obtained through use of a three-component strain gage force balance and static pressure measurement equipment at a test Reynolds number of R c = 100,000. Emphasis was placed on determining the effect of trip wire placement and size on such performance parameters as (C l /C d )max and (C l 3/2 /C d )max. Prediction of transition location by the criterion due to Tani and Gibbings was found to have limited application. Most trip wire locations resulted in degraded performance, but for some locations, minimum drag was reduced, maximum lift to drag ratio increased, and hysteresis averted.

An experimental investigation was conducted in order to document the structure and behavior of laminar separation bubbles at low Reynolds numbers. Data of this type is necessary if the currently insufficient analytical and numerical models are to be improved. The laminar separation bubble which forms on a NACA 66(3)-018 airfoil model was surveyed at chord Reynolds numbers ranging from 50,000 to 200,000 at angles of attack from 8 to 12 degrees. The effects of the various testing conditions on the separation bubble were isolated, and the data was analyzed in relation to existing separation bubble correlations in order to test their low Reynolds number applicability. This analysis indicated that the chord Reynolds number and the disturbance environment strongly influence the experimental pressure distributions. These effects must be included in any analytic prediction technique applied to the low Reynolds number flight regime.

An experimental investigation of spontaneous and forced transition on a secant-ogive-nose axisymmetric body was conducted for length Reynolds numbers between 0.315X106 and 1.03xl06. Several different transition modes were observed and documented using smoke-flow visualization and hot-wire anemometry. Sound was used to control the frequency and enhance the amplitude of the disturbances in the boundary layer during the forced transition studies. Three different mechanisms of transition originate as a viscosity-conditioned Tollmien-Schlichting (T-S) instability. The evolution of these instability waves depends upon Reynolds number and the frequency and amplitude of the disturbance. A change in the nonlinear development of the secondary instability—the A-vortex structure—could be induced by changing only the amplitude of the T-S waves, keeping all other conditions constant. The frequency spectra from a hot wire placed in this flowfield show an abrupt decrease in the psd of the subharmonic of the T-S wave frequency as the amplitude of the disturbance is increased. Mean and fluctuating velocity profiles and the mean pressure distribution along the body are also presented. The results support recent theories that describe the events in the nonlinear region.

The role of the flow visualization in providing the necessary insight for the development of theoretical models of complex afterbody and base flows is described. Methods of calculating the turbulent base pressure for axisymmetric configurations are discussed. Emphasis is placed on the treatment of supersonic flows for cylindrical, boattail, and flare bodies as well as sharp and blunt cones and base flow nozzles. The guidance provided by flow visualization in the development of theoretical models for the transonic case is also discussed. The current interests in applying finite difference techniques to these-problems indicate that the extensive use of flow visualization data will continue.

An experimental study of the effect of boundary-layer transition on the performance of the Miley airfoil at Reynolds numbers below 6xl0 was conducted. Lift and drag measurements were taken using a twocomponent strain gage force balance over a range of Reynolds numbers from 7 x 10 to 3 x 10. Static pressure distributions on the surface of the airfoil were measured for Reynolds numbers up to 6xl0 5 . Smoke flow visualization was used at Reynolds numbers between 7 x 1Q and 5 x 10 in order to obtain a qualitative picture of boundary-layer transition and flow separation. Initial studies showed a large region of hysteresis in both lift and drag performance between Reynolds numbers of 7 x 10 and 1.5 x 10. The hysteresis loop varied in size but typically occurred between 10and 18-deg angle of attack and resulted in up to a 15% difference in lift coefficient and up to a 60% difference in drag coefficient. Stability of the hysteresis loop was found to be dependent on several factors, most important of which were freestream turbulence, acoustic excitation, and boundary-layer trips. The test section environment was documented to explain changes in boundary-layer performance under normal operating conditions. Finally, quantitative analysis of the boundary-layer performance was conducted with a hot-wire anemometer at 0-, 7-, and 13-deg angle of attack for a Reynolds number of 1.5X 10. Comparison of the data obtained during the different experimental phases provides a consistent picture of the boundary-layer performance and subsequent hysteresis loop. Changes in the testing environment were found to be the critical factors in variations of the experimental results. Documentation of the test environment is a prerequisite to the analysis of any test results at low Reynolds numbers.

The results of an investigation of the influence of free stream disturbances on the lift and drag performance of the Lissaman 7769 airfoil are presented. The wind tunnel disturbance environment is described using hot-wire anemometer and sound pressure level measurements. The disturbance level is increased by the addition of a ‘turbulence screen’ upstream of the test section and/or the addition of a flow restrictor downstream of the test section. For the Lissaman airfoil it was found that the problems associated with obtaining accurate wind tunnel data at low chord Reynolds numbers (i.e., below 200,000) are compounded by the extreme sensitivity of the boundary layers to the free stream disturbance environment. The effect of free stream disturbances varies with magnitude, frequency content, and source of the disturbance.

The results from an experimental study of the boundary layer surrounding an axisymmetric model consisting of a 3-caliber secant ogive nose, a 2-caliber cylindrical midsection, and a 1-caliber 7-deg conical boattail are presented. Low-speed flow visualization data at zero angle of attack are shown for Reynolds numbers between 315,000 and 1,030,000 (based upon total body length), and ratios of peripheral to freestream velocity between 0 and 1.67. The smoke-flow photographs display the various stages of transition of the boundary layer. In addition, pressure distribution data are included for the nonspinning case. The pressure data are used in conjunction with the photographic data to present a physical picture of the boundary-layer development on the model.

The laminar separation, transition, and turbulent reattachment near the leading edge of a two-dimensional NACA 663 -018 airfoil were investigated using a low-speed, smoke visualization wind tunnel. Lift and drag force measurements were made using an external strain gage balance for a chord Reynolds number range of 40,GOO400,000. An extensive flow visualization study was performed and correlated with the force measurements. Experiments were also conducted with distributed surface roughness at the leading edge and external acoustic excitation to influence the development of the airfoil boundary layer. This study delineates the effects of angle of attack and chord Reynolds number on the separation characteristics and airfoil performance. Nomenclature c = model chord cd = section profile drag coefficient (uncorrected) cf = section lift coefficient (uncorrected) Cp = pressure coefficient / = acoustic frequency, Hz R = reattachment location Rc = Reynolds number based on chord length, U^ civ S = separation location T = location of approximate end of transition £/«, = freestream velocity x/c = nondimensional distance along chord a = angle of attack v - kinematic viscosity