Showing papers on "Airfoil published in 1983"

Noise generated by airfoil profiles placed in a uniform laminar flow

Abstract: The present paper is devoted to the experimental study of the noise generated by an airfoil profile placed in a uniform laminar flow. The far-field acoustic spectrum is shown to be composed of a broadband contribution around frequency fs and a discrete contribution at equidistant frequencies fn, which follow power laws of the forms fs ∼ U1.5 and fn ∼ U0.85. Both contributions can be accounted for by a simple model derived from the original suggestions of Tam (1974) and Fink (1975). It is essentially assumed that the diffraction of the Tollmien-Schlichting instabilities by the trailing edge generates acoustic waves which propagate in the far field and also trigger an aeroacoustic feedback loop, whose length is equal to the distance between the trailing edge and the maximum velocity point of the airfoil.

Fundamentals of Flight

Abstract: 1 Conversion Factors Between SI Units and English Units 2 Nomenclature 3 A Brief History of Aeronautics 4 The Anatomy of the Airplane 5 The Nature of Aerodynamic Forces: Dimensional Analysis 6 Theory and Experiment Wind Tunnels 7 The Atmosphere 8 Incompressible One-Dimensional Flow 9 One-Dimensional Flow in a Compressible Fluid 10 Two-Dimensional Flow: Lift and Drag 11 The Finite Wing 12 Effects of Viscosity 13 Determination of Total Incompressible Drag 14 Compressibility Drag 15 Airfoils and Wings 16 High-Lift Systems 17 Aerodynamic Performance 18 Stability and Control 19 Propulsion 20 Structures 21 Hypersonic Flow 22 Rocket Trajectories and Orbits Appendix A: Characteristics of the Standard Atmosphere (SI units) Characteristics of the Standard Atmosphere (English Units) Appendix B: Derivation of the Compressible Fluid Bernoulli Equation Appendix C: Summary of State and One-Dimensional Flow Equations Index

Analytical and Experimental Evaluation of the Heat Transfer Distribution over the Surfaces of Turbine Vanes

Abstract: Three airfoil data sets were selected for use in evaluating currently available analytical models for predicting airfoil surface heat transfer distributions in a 2-D flow field Two additional airfoils, representative of highly loaded, low solidity airfoils currently being designed, were selected for cascade testing at simulated engine conditions Some 2-D analytical methods were examined and a version of the STAN5 boundary layer code was chosen for modification The final form of the method utilized a time dependent, transonic inviscid cascade code coupled to a modified version of the STAN5 boundary layer code featuring zero order turbulence modeling The boundary layer code is structured to accommodate a full spectrum of empirical correlations addressing the coupled influences of pressure gradient, airfoil curvature, and free-stream turbulence on airfoil surface heat transfer distribution and boundary layer transitional behavior Comparison of pedictions made with the model to the data base indicates a significant improvement in predictive capability

Time-marching transonic flutter solutions including angle-of-attack effects

Abstract: Transonic aeroelastic solutions based upon the transonic small perturbation potential equation were studied. Time-marching transient solutions of plunging and pitching airfoils were analyzed using a complex exponential modal identification technique, and seven alternative integration techniques for the structural equations were evaluated. The HYTRAN2 code was used to determine transonic flutter boundaries versus Mach number and angle-of-attack for NACA 64A010 and MBB A-3 airfoils. In the code, a monotone differencing method, which eliminates leading edge expansion shocks, is used to solve the potential equation. When the effect of static pitching moment upon the angle-of-attack is included, the MBB A-3 airfoil can have multiple flutter speeds at a given Mach number.

Passive shock wave/boundary layer control for transonic airfoil drag reduction

Abstract: An investigation of the passive shock wave/boundary layer interaction control for reducing the drag in 12 percent thick circular arc and 14 percent thick supercritical airfoils was conducted in a 3 in. x 15.4 in. transonic wind tunnel at transonic Mach numbers. A porous surface with a cavity beneath it was positioned on the area of the airfoils, mounted on the test section bottom wall, where the shock wave occurs. The static pressure distributions over the airfoil, the wake impact pressure data for determining the profile drag, and the Schlieren photographs for porous surface airfoils are presented and compared with the results for solid surface airfoils. With the porous surface the normal shock wave for solid surface was changed to a lambda shock wave system, and the wake impact pressure data indicated an appreciable drag reduction for both airfoils with the porous surface at transonic speeds while causing little or no loss of lift. The effect of porosity and cavity size is investigated and off-design performance is discussed.

The Influence of Free-Stream Disturbances on Low Reynolds Number

Abstract: 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 com- pounded 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.

Development of Advanced Circulation Control Wing High Lift Airfoils

Abstract: : Recent experimental and flight test programs have developed and confirmed the high lift capability of the Circulation Control Wing (CCW) concept. These CCW airfoils employ tangential blowing of engine bleed air over circular or near circular trailing edges, and are capable of usable lift coefficients triple those of simple mechanical flaps. Earlier versions of these blown airfoils made use of relatively complex leading and trailing edge devices which would have to be retracted mechanically for cruise flight. In a continuing program to reduce the complexity, size and weight of the CCW system, several series of advanced CCW airfoils have been developed which can provide STOL capability for both military and commercial aircraft using much smaller, less complex high lift systems. This paper describes these configurations and presents the experimental results confirming their aerodynamic characteristics. Comparisons to previous CCW and more conventional high lift systems are provided.

The influence of free-stream disturbances on low Reynolds number airfoil experiments

Abstract: 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 effect of a leading-edge slat on the dynamic stall of an oscillating airfoil

Abstract: The dynamic stall characteristics of a slatted airfoil were investigated experimentally on a 2-ft-chord airfoil oscillating in pitch at M = 0.2 for a range of reduced frequency and mean angle of oscillation. The slat produced a flow that remained attached to the airfoil for angles well above those normally attained by the retreating blade of a helicopter during high speed flight. The dynamic stall vortex usually associated with these flight conditions was completely eliminated for all angles under 30 deg. Instantaneous surface pressure, lift, and pitching moment data are presented as a function of incidence throughout the oscillation cycle; a detailed analysis of instantaneous boundary-layer flow behavior for the various test conditions is included.

Friction damping of flutter in gas turbine engine airfoils

Abstract: This paper investigates the feasibility of using blade-to-ground friction dampers to stabilize flutter in blades. The response of an equivalent one mode model in which the aerodynamic force is represented as negative viscous damping is examined to investigate the following issues: the range of amplitudes over which friction damping can stabilize the response, the maximum negative aerodynamic damping that can be stabilized in such a manner, the effect of simultaneous resonant excitation on these stability limits, and the determination of those damper parameters which will be the best for flutter control.

The lift on an aerofoil in starting flow

Abstract: An analysis is given of the initial development of the lift on an aerofoil in inviscid starting flow. It is shown that because of the spiral shape of the vortex sheet shed initially from the trailing edge the lift and drag are both singular at the start of impulsive motion. This result is in contrast with the prediction of finite forces by methods that assume the vortex sheet to be initially planar. The effect of a steady rate of change of incidence following the sudden onset of transverse (heaving) motion of an aerofoil in a steady stream is also discussed.

Dynamic-stall regulation of the Darrieus turbine

Abstract: A two-dimensional unsteady airfoil analysis is described which utilizes a doublet panel method to model the airfoil surface, an integral boundary scheme to model the viscous attached flow, and discrete vortices to model the detached boundary layers which form the airfoil wake region. This model has successfully predicted steady lift and drag coefficients as well as pressure distributions for several airfoils with both attached and detached boundary layers. Unsteady calculations have thus far been limited to attached flow situations. Instantaneous pressure distributions have also been obtained on a single-bladed rotor operating in a tow tank in order to provide experimental data for eventual comparison with analytical predictions.

Self-adjusting wind power machine

Abstract: A machine for economical recovery of wind power employs a self-adjusting mass-balanced aerodynamic blade weathervaning freely around a lengthwise pitching axis forward of its aerodynamic center, and an aerodynamic roller in its leading edge, spun at high RPM by a motor. The roller controls aerodynamic performance to high levels of efficiency at high lift coefficients, employing novel roller/airfoil profiles. For ship propulsion, the self-adjusting blade with roller stopped is like a furled sail, and with the blade held angling to the wind with roller spinning is like a large, efficient, easily controlled sail. On a horizontal axis wind turbine, the self-adjusting blade is continuously held to an efficient angle of attack by centrifugal lift-increasing pitching moments balancing aerodynamic lift-decreasing pitching moments. The blade whirls steadily despite fluctuations of wind speed and direction, reducing stresses, and preventing structural damage or loss of efficiency. Gyroscopic blade teetering moments are neutralized by mass-balance counter-spin, using a single blade with a balancing stub, on a teetering pivot at the mill shaft. A teetering pivot at the top of the mast and horizontal tail surfaces complete an overall dynamic stress relief system. Sensors monitor operating conditions, control roller speed and move centrifugal mass positions for optimum power output.

Wind turbine apparatus

Abstract: Wind turbine apparatus includes a plurality of air foil sail elements secured to a circular frame rotatable in response to wind reacting with the sail elements. The sail elements include deformable outer skin portions, one of which is flattened against an interior form in response to wind forces, and the other of which extends convexly away from the interior form. The deformation changes the camber of the sail element.

Numerical simulation of airfoil ice accretion

Abstract: A mathematical model of glaze and rime ice accretion on a two-dimensional airfoil is presented. The model employs standard methods for calculating the flow field and cloud droplet trajectories. This data is used as input to a thermodynamic analysis of the ice accretion process which includes liquid runback. The structure of the model allows for dynamic updating of the droplet collection efficiency and heat and mass transfer processes as the accreted ice shape changes. This results in the improved shape prediction over previously described methods. Simulations made with the model are compared to experimental results obtained for a NACA 0012 airfoil in the NASA/Lewis Icing Research Tunnel. The agreement with experiment is found to be generally satisfactory for some cases, but a need for improvements is indicated by others. Suggestions for improvements are made.

Interactions of Airfoils with Gusts and Concentrated Vortices in Unsteady Transonic Flow

Abstract: Unsteady interactions of concentrated vortices and distributed free-stream gusts with a stationary airfoil have been analyzed in two-dimensional transonic flow. A simple method of introducing such disturbances has been implemented numerically in the well-known transonic small-disturbance code LTRAN2, and calculations have been performed for two important classes of current aerodynamic problems. The first, which demonstrates many of the essential features of the interactions between helicopter rotor blades and their trailing-vortex wakes, is that of a discrete potential vortex convecting past an airfoil. The second is the response of a transonic airfoil to a transverse periodic gust, with and without the alleviation that can be achieved by the proper active control motion of a trailing-edge flap. In both cases, unsteady effects are found to play important roles in the shock-wave motion, in the overall flow-field development, and consequently, in the air loads on the airfoil.

Passive drag control of airfoils at transonic speeds

Abstract: An airfoil for transonic speeds includes a porous top surface extending from a location about 50 to 60% of the chord length from a leading edge of the airfoil to a location about 80 to 90% of the chord length from the leading edge. A cavity is defined under the porous surface in the airfoil which has a depth of from 0.05 to 0.2% of the chord length. The porosity of the porous surface is chosen to be from 1 to 3% of the total airfoil surface and may be variable. The presence of the porous surface and cavity decrease airfoil drag at transonic speeds by providing a pathway between a high pressure area downstream of a shock wave formed on the airfoil at transonic speeds to a low pressure area within a bubble on the airfoil upstream of the shock wave.

Rotor-vortex interaction noise

Abstract: A theoretical and experimental study was conducted to develop a validated first principles analysis for predicting noise generated by helicopter main-rotor shed vortices interacting with the tail rotor. The generalized prediction procedure requires a knowledge of the incident vortex velocity field, rotor geometry, and rotor operating conditions. The analysis includes compressibility effects, chordwise and spanwise noncompactness, and treats oblique intersections with the blade planform. Assessment of the theory involved conducting a model rotor experiment which isolated the blade-vortex interaction noise from other rotor noise mechanisms. An isolated tip vortex, generated by an upstream semispan airfoil, was convected into the model tail rotor. Acoustic spectra, pressure signatures, and directivity were measured. Since assessment of the acoustic prediction required a knowledge of the vortex properties, blade-vortes intersection angle, intersection station, vortex stength, and vortex core radius were documented. Ingestion of the vortex by the rotor was experimentally observed to generate harmonic noise and impulsive waveforms.

Linear/nonlinear behavior in unsteady transonic aerodynamics

Abstract: The accurate calculation of the aerodynamic forces in unsteady transonic flow requires the solution of the nonlinear flow equations. The aeroelastician, on the other hand, seeks to treat his problems (flutter, for example) by means of linear equations whenever possible. He may do this, even when the underlying flow is nonlinear, if the perturbation forces are linear over some (perhaps small) range of unsteady amplitude of motion. This paper assesses the range of parameters over which linear behavior occurs. In particular calculations are made for an NACA 64A006 airfoil oscillating in pitch over a range of amplitudes, frequencies, and Mach numbers. The primary aerodynamic method used is the well known LTRAN2 code of Ballhaus and Goorjian that provides a finite-difference solution to the low frequency, small disturbance, two-dimensional potential flow equation. Comparisons are made with linear subsonic theory, local linearization, and, for steady flow, with the full potential equation code of Bauer, Garabedian, and Korn.

Analysis of airfoil leading-edge separation bubbles

Abstract: A local inviscid-viscous interaction technique was developed for the analysis of low speed airfoil leading edge transitional separation bubbles. In this analysis an inverse boundary layer finite difference analysis is solved iteratively with a Cauchy integral representation of the inviscid flow which is assumed to be a linear perturbation to a known global viscous airfoil analysis. Favorable comparisons with data indicate the overall validity of the present localized interaction approach. In addition numerical tests were performed to test the sensitivity of the computed results to the mesh size, limits on the Cauchy integral, and the location of the transition region.

Analysis of unswept and swept wing chordwise pressure data from an oscillating NACA 0012 airfoil experiment. Volume 1: Technical Report

Abstract: The unsteady chordwise force response on the airfoil surface was investigated and its sensitivity to the various system parameters was examined. A further examination of unsteady aerodynamic data on a tunnel spanning wing (both swept and unswept), obtained in a wind tunnel, was performed. The main body of this data analysis was carried out by analyzing the propagation speed of pressure disturbances along the chord and by studying the behavior of the unsteady part of the chordwise pressure distribution at various points of the airfoil pitching cycle. It was found that Mach number effects dominate the approach to and the inception of both static and dynamic stall. The stall angle decreases as the Mach number increases. However, sweep dominates the load behavior within the stall regime. Large phase differences between unswept and swept responses, that do not exist at low lift coefficient, appear once the stall boundary is penetrated. It was also found that reduced frequency is not a reliable indicator of the unsteady aerodynamic response in the high angle of attack regime.

An exploratory study of finite difference grids for transonic unsteady aerodynamics

Abstract: Unsteady aerodynamic forces are calculated by the XTRAN2L finite difference program which solves the complete two dimensional unsteady transonic small perturbation equation. The unsteady forces are obtained using a pulse transfer function technique which assumes the flow field behaves in a locally linear fashion about a mean condition. Forces are calculated for a linear flat plate using the default grids from the LTRAN2-NLR, LTRAN2-HI, and XTRAN3S programs. The forces are compared to the exact theoretical values for flat plate, and grid generated boundary and internal numerical reflections are observed to cause significant errors in the unsteady airloads. Grids are presented that alleviate the reflections while reducing computational time up to fifty-three percent and program size up to twenty-eight percent. Forces are presented for a six percent thick parabolic arc airfoil which demonstrate that the transform technique may be successfully applied to nonlinear transonic flows.

Synthesized airfoil data method for prediction of dynamic stall and unsteady airloads

Abstract: A detailed analysis of dynamic stall experiments has led to a set of relatively compact analytical expressions, called synthesized unsteady airfoil data, which accurately describe in the time-domain the unsteady aerodynamic characteristics of stalled airfoils. An analytical research program was conducted to expand and improve this synthesized unsteady airfoil data method using additional available sets of unsteady airfoil data. The primary objectives were to reduce these data to synthesized form for use in rotor airload prediction analyses and to generalize the results. Unsteady drag data were synthesized which provided the basis for successful expansion of the formulation to include computation of the unsteady pressure drag of airfoils and rotor blades. Also, an improved prediction model for airfoil flow reattachment was incorporated in the method. Application of this improved unsteady aerodynamics model has resulted in an improved correlation between analytic predictions and measured full scale helicopter blade loads and stress data.

Device for raising and lowering an airfoil

Abstract: A device for raising and lowering an airfoil mounted on a lower portion of a vehicle body comprises a driving mechanism for driving the airfoil upwardly and downwardly, an electric device such as a motor for actuating the driving mechanism, a vehicle height detector for detecting the height of the vehicle body with respect to an axle, an airfoil height detector for detecting the height of the airfoil with respect to the vehicle body, and a control unit responsive to the vehicle height as detected by the vehicle height detector, the airfoil height as detected by the airfoil height detector, and the speed of travel of the vehicle body for actuating the electric device to lower the airfoil when the speed exceeds a predetermined level and the vehicle height is high.

Comparison of NACA 6-series and 4-digit airfoils for Darrieus wind turbines

Abstract: The aerodynamic efficiency of Darrieus wind turbines as effected by blade airfoil geometry was investigated. Analysis was limited to curved-bladed machines having rotor solidities of 7 to 21% and operating at a Reynolds number of 3 X 10/sup 6/, Ten different airfoils, having thickness-to-chord ratios of 12, 15, and 18%, were studied. Performance estimates were made using a blade element/momentum theory approach. Results indicated that NACA 6-series airfoils yeild peak power coefficients as great as NACA 4-digit airfoils and have broader and flatter power coefficient-tip speed ratio curves. Sample calculations for an NACA 63/sub 2/-015 airfoil showed an annual energy output increase of 17-27%, depending on rotor solidity, compared to an NACA 0015 airfoil.

Boundary layer characteristics of the Miley airfoil at low Reynolds numbers

Abstract: 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.

Unsteady Aerodynamics and Gapwise Periodicity of Oscillating Cascaded Airfoils

Abstract: Tests were conducted on a linear cascade of airfoils oscillating in pitch to measure the unsteady pressure response on selected blades along the leading edge plane of the cascade and over the chord of the center blade. The pressure data were reduced to Fourier coefficient form for direct comparison and were also processed to yield integrated loads and, particularly, the aerodynamic damping coefficient. In addition, results from two unsteady theories for cascaded blades with nonzero thickness and camber were compared with the experimental measurements. The three primary results that emerged from this investigation were: (a) from the leading edge plane blade data, the cascade was judged to be periodic in unsteady flow over the range of parameters tested, (b) as before, the interblade phase angle was found to be the single most important parameter affecting the stability of the oscillating cascade blades, and (c) the real blade theory and the experiment were in excellent agreement for the several cases chosen for comparison.

Wing tip airfoils

Abstract: The subject invention relates to a new and improved airfoil design. More particularly, an airfoil design is disclosed which improves efficiency by taking advantage of the vortex created at the wing tips.