Showing papers on "Airfoil published in 1982"
07 Jun 1982
777 citations
TL;DR: In this paper, 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.
Abstract: 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
240 citations
01 Sep 1982
TL;DR: In this article, a method for calculating the output power from large horizontal-axis wind turbines is presented, where modifications to the airfoil characteristics and the momentum portion of classical blade element-momentum theory are given that improve correlation with measured data.
Abstract: A method for calculating the output power from large horizontal-axis wind turbines is presented. Modifications to the airfoil characteristics and the momentum portion of classical blade element-momentum theory are given that improve correlation with measured data. Improvement is particularly evident at low tip-speed ratios where aerodynamic stall can occur as the blade experiences high angles of attack. Output power calculated using the modified theory is compared with measured data for several large wind turbines. These wind turbines range in size from the DOE/NASA 100 kW Mod-0 (38 m rotor diameter) to the 2000 kW Mod-1 (61 m rotor diameter). The calculated results are in good agreement with measured data from these machines.
182 citations
01 Feb 1982
136 citations
01 Dec 1982
TL;DR: In this article, the authors describe the techniques developed for analysis and evaluation of the hot film and hot wire signals, offer some interpretation of the results, and tabulate all the cases in which flow reversal has been recorded.
Abstract: Detailed unsteady boundary layer measurements are presented for eight airfoils oscillated in pitch through the dynamic stall regime. The present report (the third of three volumes) describes the techniques developed for analysis and evaluation of the hot film and hot wire signals, offers some interpretation of the results, and tabulates all the cases in which flow reversal has been recorded.
134 citations
Patent•
30 Dec 1982
TL;DR: In this paper, a variable camber mechanism for the leading and/or trailing edge of an airfoil that is operable for lift variation at high airspeeds of a jet aircraft such as for maneuvering of fighter aircraft is presented.
Abstract: A variable camber apparatus for the leading and/or trailing edge of an airfoil that is operable for lift variation at high airspeeds of a jet aircraft such as for maneuvering of fighter aircraft. The cambering apparatus comprises an upper surface flexible skin panel which is supported along its inner edge by a wing spar assembly and along its outermost edge, by an airfoil edge forming structure. A hinged rib member, through a kinematic linkage mechanism associated between itself and the structure at the outermost edge of the airfoil section, bends and torsionally twists the airfoil edge forming structure about a relative spanwise axis. The kinematic linkage mechanism is slaved to rotating means of the rib member, for torsionally twisting the outermost edge of the airfoil section, upper surface portion, the precise amount of rotation so as to flexuously bend the upper surface and thereby contour it to conform to an aerodynamically predetermined curvilinear plot that will produce the desired camber and the change in the lift characteristics of the airfoil surface.
128 citations
01 Jul 1982
TL;DR: In this paper, the static and dynamic characteristics of seven helicopter sections and a fixed-wing supercritical airfoil were investigated over a wide range of nominally two dimensional flow conditions, at Mach numbers up to 0.30 and Reynolds number up to 4 x 10 to the 6th power.
Abstract: The static and dynamic characteristics of seven helicopter sections and a fixed-wing supercritical airfoil were investigated over a wide range of nominally two dimensional flow conditions, at Mach numbers up to 0.30 and Reynolds numbers up to 4 x 10 to the 6th power. Details of the experiment, estimates of measurement accuracy, and test conditions are described in this volume (the first of three volumes). Representative results are also presented and comparisons are made with data from other sources. The complete results for pressure distributions, forces, pitching moments, and boundary-layer separation and reattachment characteristics are available in graphical form in volumes 2 and 3. The results of the experiment show important differences between airfoils, which would otherwise tend to be masked by differences in wind tunnels, particularly in steady cases. All of the airfoils tested provide significant advantages over the conventional NACA 0012 profile. In general, however, the parameters of the unsteady motion appear to be more important than airfoil shape in determining the dynamic-stall airloads.
117 citations
01 Sep 1982
TL;DR: In this article, experimentally derived force and moment data are presented for eight airfoil sections that were tested at fixed and varying incidence in a subsonic two dimensional stream.
Abstract: Experimentally derived force and moment data are presented for eight airfoil sections that were tested at fixed and varying incidence in a subsonic two dimensional stream. Airfoil incidence was varied through sinusoidal oscillations in pitch over a wide range of amplitude and frequency. The surface pressure distribution, as well as the lift, drag, and pitching moment derived therefrom, are displayed in a uniform fashion to delineate the static and dynamic characteristics of each airfoil both in and out of stall.
117 citations
TL;DR: In this paper, it was shown that the transonic potential flow partial differential equation admits nonsymmetric solutions with large positive or negative lift, for symmetric airfoils at zero angle of attack.
Abstract: The two-dimensional transonic potential flow equation, when solved in discrete form for steady flow over an airfoil, has been found to yield more than one solution in certain bands of angle of attack and Mach number. The most striking ex- ample of this is the appearance of nonsymmetric solutions with large positive or negative lift, for symmetric airfoils at zero angle of attack. The behavior of these "anomalous" solutions is exam- ined as grid size is varied by large factors and found to be not qualitatively different from that of %ormalrf solutions (outside the nonuniqueness band). Thus it appears that the effect is not due to discretization error, and that the basic tran- sonic potential flow partial differential equation admits nonunique solutions for certain values of angle of attack and Mach number.
97 citations
TL;DR: In this article, a transonic flow over a supercritical swept wing is considered, and a wave-drag expression is derived for the conservation form of the momentum equation, expressed in terms of a small-disturbance equation.
Abstract: Computational fluid dynamics are used to discuss problems inherent to transonic three-dimensional flow past supercritical swept wings. The formulation for a boundary value problem for the flow past the wing is provided, including consideration of weak shock waves and the use of parabolic coordinates. A swept wing code is developed which requires a mesh of 152 x 10 x 12 points and 200 time cycles. A formula for wave drag is calculated, based on the idea that the conservation form of the momentum equation becomes an entropy inequality measuring the drag, expressible in terms of a small-disturbance equation for a potential function in two dimensions. The entropy inequality has been incorporated in a two-dimensional code for the analysis of transonic flow over airfoils. A method of artificial viscosity is explored for optimum pressure distributions with design, and involves a free boundary problem considering speed over only a portion of the wing.
85 citations
TL;DR: In this article, the authors investigated the ability of a free vortex by a Joukowski airfoil in its vicinity for lift augmentation through a two-dimensional potential flow analysis.
Abstract: The ability of trapping a free vortex by a Joukowski airfoil in its vicinity for lift augmentation is investigated through a two-dimensional potential flow analysis. The effects of angle of attack, thickness, and camber on the trapping performance of the airfoil are examined separately. It is found that to capture a vortex which is stable to small disturbances, the proper arrangement is to have a rather thick symmetric airfoil at a small angle of attack.
TL;DR: In this paper, an experimental investigation of the near wake of a thin airfoil at various incidence angles was performed in a wind tunnel, and the wake structure was measured using hot-wire sensors.
Abstract: An experimental investigation of the near wake of a thin airfoil at various incidence angles is reported in this paper. The airfoil (NACA 0012 basic thickness form) was located in a wind tunnel, and the wake structure was measured using hot-wire sensors. The measurements of mean-velocity, turbulence intensity and Reynolds-stress components across the wake at several distances downstream show the complex nature of the near wake and its asymmetrical behavior. The asymmetry in the wake property, which is maintained up to a length of 1.5 chords downstream of the trailing edge of the blade, is dependent on the incidence angle of the inlet flow. The streamwise velocity defect in an asymmetric wake decays more slowly compared to that of a symmetric wake. The streamline curvature due to the blade loading has a substantial effect on the mean velocity profile as well as the turbulence structure. The numerical study of the same wake indicates that the existing turbulence closure models need some modification to account for the asymmetric characteristics of the wake.
TL;DR: In this article, a wing of 8 in. span and 2 in. chord was oscillated in a low-speed wind tunnel and the average thrusting effort of the wing was measured and plotted in coefficient form against reduced frequency, with pitching amplitude and phase angle as parameters.
Abstract: In this experimental study of flapping-wing thrust, a wing of 8 in. span and 2 in. chord was oscillated in a lowspeed wind tunnel. The driving apparatus produced nearly sinusoidal heaving with superimposed pitching of variable amplitude and phase angle. The flapping frequency range of 0-8 Hz produced reduced frequencies for which flow separation was unlikely to occur, based on the measured static characteristics of the test airfoil (NACA 0012) over the Reynolds number range of interest (25,000-40,000). The average thrusting effort of the wing was measured and plotted in coefficient form against reduced frequency, with pitching amplitude and phase angle as parameters. Comparisons of the results with theoretical predictions and previous experimental work were made. In general, the results show approximately linear dependence of thrust on reduced frequency and best performance at phase angles of 90-120 deg of pitching lagging heaving. Although thrusting effort was produced for all pitching amplitudes, including zero, the highest readings were obtained for the maximum pitching amplitude of 12.1 deg.
TL;DR: In this article, the aeroelastic stability of a helicopter rotor blade in hover is examined using a finite element formulation based on the principle of virtual work, where the rotor blade is discretized into beam elements, each with ten modal degrees of freedom.
Abstract: The aeroelastic stability of flap bending, lead-lag bending, and torsion of a helicopter rotor blade in hover is examined using a finite element formulation based on the principle of virtual work. Quasi-steady two-dimensional airfoil theory is used to obtain the aerodynamic loads. The rotor blade is discretized into beam elements, each with ten modal degrees of freedom. The resulting nonlinear equations of motion are solved for steady-state blade deflections through an iterative procedure. The flutter solution is calculated assuming blade motion to be a small perturbation about the steady solution. The normal mode method based on the coupled rotating natural modes about the steady deflections is used to reduce the number of equations in the flutter eigenanalysis. Results are presented for hingeless and articulated rotor blade configurations.
TL;DR: In this paper, the aerofoil stall of streamlined flow near the rounded leading edge of an aerodynamic body was studied theoretically for large Reynolds number motion, where the small angle of attack is raised above its critical stall value, resulting in a very pronounced local bulge appearing in the flow displacement.
Abstract: The unsteady breakdown or stall of streamlined flow near the rounded leading edge of an aerofoil, as the small angle of attack is raised above its critical stall value, is studied theoretically for large Reynolds number motion. The unsteady developments take place first over a relatively slow time scale but then the corresponding solution breaks down with a singularity, forcing a switch to a faster and more nonlinear process. The latter involves a very pronounced local bulge appearing in the flow displacement, accompanied by reversed flow at the aerofoil surface, and comparisons with experimental observations of dynamic stall are noted.
TL;DR: In this article, an asymptotic theory of the Navier-Stokes equations at large Reynolds numbers is presented, where the boundary value problem is reduced to an integrodifferential equation for the distribution of the friction.
Abstract: The two-dimensional flow of a viscous incompressible fluid near the leading edge of a slender airfoil is considered. An asymptotic theory of this flow is constructed on the basis of an analysis of the Navier—Stokes equations at large Reynolds numbers by means of matched asymptotic expansions. A central feature of the theory is the region of interaction of the boundary layer and the exterior inviscid flow; such a region appears on the surface of the airfoil in a definite range of angles of attack. The boundary-value problem for this region is reduced to an integrodifferential equation for the distribution of the friction. This equation has been solved numerically. As a result, closed separation regions are constructed, and the angle of attack at which separation occurs is found.
01 Jan 1982
TL;DR: The results of an experimental study to document the effects of separation and transition on the performance of an airfoil designed for low Reynolds number operation are presented in this paper, where lift, drag, and flow visualization data were obtained for the Eppler 61 airfoIL section for chord Reynolds numbers from about 30,000 to over 200,000.
Abstract: The results of an experimental study to document the effects of separation and transition on the performance of an airfoil designed for low Reynolds number operation are presented. Lift, drag and flow visualization data were obtained for the Eppler 61 airfoil section for chord Reynolds numbers from about 30,000 to over 200,000. Smoke flow visualization was employed to document the boundary layer behavior and was correlated with the Eppler airfoil design and analysis computer program. Laminar separation, transition and turbulent reattachment had significant effects on the performance of this airfoil.
TL;DR: The results of a study supported by NASA under the Energy Efficient Engine Program, conducted to investigate the development of boundary layers under the influence of velocity distributions that simulate the suction sides of two state-of-the-art turbine airfoils, are presented in this paper.
Abstract: The results of a study supported by NASA under the Energy Efficient Engine Program, conducted to investigate the development of boundary layers under the influence of velocity distributions that simulate the suction sides of two state-of-the-art turbine airfoils, are presented. One velocity distribution represented a forward loaded airfoil ('squared-off' design), while the other represented an aft loaded airfoil ('aft loaded' design). These velocity distributions were simulated in a low-speed, high-aspect-ratio wind tunnel specifically designed for boundary layer investigations. It is intended that the detailed data presented in this paper be used to develop improved turbulence model suitable for application to turbine airfoil design.
TL;DR: In this article, the effect of thickness increases with frequency, with thick airfoils being quieter than thin ones, and it is found that the effect is large and must be accounted for in any fundamental airfoil noise theory that attempts to describe the noise emitted from real airfoILS.
Abstract: Noise emission from very small chord and very large chord airfoils was measured with eleven 0.63 cm microphones placed along a horizontal semicircle (4.57 m radius) that was centered at the leading edge of the test airfoil. The noise signals were analyzed by an automated spectrum analyzer which yielded 1/3-octave band sound pressure level spectra for each microphone, and the data were corrected to remove the effects of atmospheric attenuation and jet noise. It is found that the effect of thickness is large and must be accounted for in any fundamental airfoil noise theory that attempts to describe the noise emitted from real airfoils. Incident mean velocity gradients and compressibility must also be taken into account. The effect of thickness increases with frequency, with thick airfoils being quieter than thin ones.
01 Apr 1982
TL;DR: A wind tunnel model mount system for conducting flutter research using a rigid wing was developed in this paper, where the wing is attached to a splitter plate so that the two move as one rigid body.
Abstract: A wind tunnel model mount system for conducting flutter research using a rigid wing was developed The wing is attached to a splitter plate so that the two move as one rigid body The splitter plate is supported away from the tunnel wall by a system of rods with fixed fixed and conditions The rods flex in such a way that only pitch and plunge oscillations are permitted At the tunnel wall the rods are attached to a remotely controlled turntable so that angle of attack can be varied Wind tunnel data obtained by using the mount system are presented for a supercritical and a conventional airfoil Both classical flutter and stall flutter data are presented
01 Jan 1982
TL;DR: In this paper, a transonic small perturbation potential equation was used to determine transonic flutter boundaries versus Mach number and angle of attack for NACA 64A010 and MBB A-3 airfoils.
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.
TL;DR: In this paper, the authors used the apparent masses to compute the initial lift and drag of an airfoil that starts impulsively from rest in viscid incompressible flow past a slightly cambered airframe at a small angle of attack.
Abstract: The method of apparent masses is utilized to compute the initial lift and drag of an airfoil that starts impulsively from rest. Analytical solutions are obtained for in viscid incompressible flow past a slightly cambered airfoil at a small angle of attack. For a Joukowski airfoil with a cusped trailing edge, it is found that increasing camber or angle of attack will cause increases in both initial lift and drag, whereas increasing thickness will result in an opposite effect. Effects of trailing-edge angle are examined by considering the symmetric Karman-Trefftz airfoil. The result shows that both lift and drag vanish at the initial instant if the airfoil has a finite trailing-edge angle.
TL;DR: In this paper, the authors simulate a retreating-blade stall with simultaneous fluctuations of both velocity and incidence on an airfoil executing oblique fore-and-aft translations and show that when dynamic stall occurs in out-of-phase or in-phase variations of incidence and velocity, strong unsteady effects are present.
Abstract: Proper simulation of a retreating-blade stall requires considering incidence fluctuations combined withsimultaneous velocity oscillations of the incoming airstream. This paper reports experimental results obtained when simultaneous fluctuations of both velocity and incidence are generated on an airfoil executing oblique fore-and-aft translations. When dynamic stall occurs in out-of-phase or in-phase variations of incidence and velocity, strong unsteady effects are present. In the first case, an instantaneous lift overshoot and a time delay between incidences of dynamic and static stall are observed. In the second case, the main effect consists of an overshoot of the mean aerodynamic coefficients over the period, owing to a cyclic separation and reattachment process with a strong vortex shedding all along the upper surface during the separation phase.
TL;DR: In this paper, the effect of Mach number on transonic flutter speed was studied for several values of four different aeroelastic parameters, including angle of attack, angle of freedom, and CAST 7.
Abstract: Transonic flutter analyses are performed for two conventional airfoils, NACA 64A006 and NACA 64A010, and three supercritical airfoils, MBB A-3, CAST 7, and the NASA TF-8A wing section at 65.3% semispan. Two degrees of freedom, plunging and pitching about the quarter-chord axis, are considered. The aerodynamic data are obtained by using the two transonic aerodynamics codes LTRAN2 (indicial and time integration methods) and STRANS2/UTRANS2 (harmonic analysis method). The unsteady aerodynamic data are computed within the low reduced frequency range. For all airfoils, the effect of Mach number on flutter speed is studied for several values of four different aeroelastic parameters. For the MBB A-3 supercritical airfoil, the effect of angle of attack on flutter speed is studied. For the cases of a flat plate and a NACA 64A006 airfoil, time response results are obtained by LTRAN2. Applicability and limitations of the two transonic codes are evaluated. Results for the transonic flutter characteristics of these airfoils are discussed and some comparisons are made. ah b,c
01 Mar 1982
TL;DR: A large chord swept supercritical LFC airfoil has been constructed for NASA-Langley's research program to determine the compatibility of supercritical airfoils with suction laminarization and to establish a technology base for future transport designs as discussed by the authors.
Abstract: A large chord swept supercritical LFC airfoil has been constructed for NASA-Langley's research program to determine the compatibility of supercritical airfoils with suction laminarization and to establish a technology base for future transport designs. Features include a high design Mach number and shock-free flow, as well as the minimization of the laminarization suction through a choice of airfoil geometry and pressure distribution. Two suction surface concepts and a variety of hybrid suction concepts involving combinations of natural and forced laminar flow are to be investigated. The test facility has been modified to insure achievement of required flow quality and transonic interference-free flow over the yawed LFC airfoil.
01 Dec 1982
Patent•
25 Jun 1982
TL;DR: In this article, an airfoil which has particular application to the blade or blades of rotor aircraft such as helicopters and aircraft propellers is designed to maintain a near zero pitching moment coefficient over a wide range of lift coefficients and to increase the drag divergence Mach number resulting in superior aircraft performance.
Abstract: This invention is an airfoil which has particular application to the blade or blades of rotor aircraft such as helicopters and aircraft propellers. The airfoil thickness distribution and camber are shaped to maintain a near zero pitching moment coefficient over a wide range of lift coefficients and provide a zero pitching moment coefficient at section Mach numbers near 0.80 and to increase the drag divergence Mach number resulting in superior aircraft performance.
TL;DR: In this article, a numerical scheme for the solution of two-dimensional time-dependent Reynolds equations is developed and applied to the flow past airfoils, where the airfoil is mapped conformally into a circular cylinder, and the equations are solved in the cylinder plane.
Abstract: A numerical scheme for the solution of two-dimensional time-dependent Reynolds equations is developed and applied to the flow past airfoils. The airfoil is mapped conformally into a circular cylinder, and the equations are solved in the cylinder plane. The turbulence models used in this study are the mixing-length model and a hybrid model consisting of mixing-length model and the two-equation k-€ model. An explicit integral relation is used to determine the kinematics of the problem. After the method was calibrated using simpler geometries, it was applied to the case of a 12% thick Joukowski airfoil at 15 deg angle of attack, at a Reynolds number of 3.6 X10. It is observed from the results that the turbulent flow past an airfoil at high angles of attack shows oscillatory behavior (with the lift and drag coefficients continuously varying with time) similar to the behavior in laminar flows.
TL;DR: In this article, the effects of various kinds of aeroelastic parameters on flutter speeds for the bending-torsion, bendingaileron, and torsion-ailerons branches are studied.
Abstract: Flutter and time-response analyses are performed for a NACA 64A006 conventional and a MBB A-3 supercritical airfoil, both oscillating with plunge, pitch, and aileron pitch degrees-of-freedom (DOF's) in smalldisturbance transonic flow. The aerodynamic coefficients are calculated using the transonic code LTRAN2NLR. The effects of various kinds of aeroelastic parameters on flutter speeds for the bending-torsion, bendingaileron, and torsion-aileron branches are studied. The flutter speeds associated with the bending-torsion branch are plotted against Mach number for different parameter values and the transonic dip phenomenon is demonstrated. To study the flutter modes, the flutter speed, amplitude ratio, and phase difference at different Mach numbers are plotted against the mass ratio for both a 2DOF and a 3DOF case. Time-response results are obtained for the NACA 64A006 and the MBB A-3 airfoils at M=Q.85 and 0.765, respectively. Based on the same sets of parameter values, the flight speeds used to obtain all the neutrally stable responses are very close to the flutter speeds obtained in the flutter analysis. The principle of linear superposition of airloads is used in the flutter analysis, but not in the response analysis.
01 Jan 1982
TL;DR: In this paper, the LEWICE program was formulated to solve a set of equations which describe the physical processes which occur during accretion of ice on an airfoil, including heat transfer in a time dependent mode, with the restriction that the flow must be describable by a 2D flow code.
Abstract: The progress toward development of a computer model suitable for predicting icing behavior on airfoils over a wide range of environmental conditions and airfoils shapes is reported. The LEWICE program was formulated to solve a set of equations which describe the physical processes which occur during accretion of ice on an airfoil, including heat transfer in a time dependent mode, with the restriction that the flow must be describable by a two-dimensional flow code. Input data comprises the cloud liquid water content, mean droplet diameter, ambient air temperature, air velocity, and relative humidity. A potential flowfield around the airfoil is calculated, along with the droplet trajectories within the flowfield, followed by local values of water droplet collection efficiency at the impact points. Both glaze and rime ice conditions are reproduced, and comparisons with test results on icing of circular cylinders showed good agreement with the physical situation.