Showing papers on "Airfoil published in 1973"

A Class of Airfoils Designed for High Lift in Incompressible Flow

Abstract: The problem studied is that of designing a single element airfoil which provides the maximum possible lift in an unseparated incompressible flow. First, an airfoil velocity distribution is defined and optimized using boundary-layer theory and the calculus of variations. The resulting velocity distribution is then used as an input for an inverse airfoil design program which provides the corresponding airfoil shape. Since there is no guarantee that an arbitrarily defined velocity distribution will yield a physically possible airfoil shape, some parametric adjustments in the optimized distributions are required in order to obtain realistic and practical airfoil geometries. Wind-tunnel tests of two different airfoils (one assuming a laminar rooftop and the other a turbulent rooftop) have been conducted and in both cases the results met the theoretically predicted performance; for example, the laminar section exhibited a low drag range of CD — 0.0085 from CL - 0.8 to CL = 2.2.

Low speed aerodynamic characteristics of a 17 percent thick airfoil section designed for general aviation applications

Abstract: Wind-tunnel tests have been conducted to determine the low-speed two-dimensional aerodynamic characteristics of a 17-percent-thick airfoil designed for general aviation applications (GA(W)-1). The results were compared with predictions based on a theoretical method for calculating the viscous flow about the airfoil. The tests were conducted over a Mach number range from 0.10 to 0.28. Reynolds numbers based on airfoil chord varied from 2.0 million to 20.0 million. Maximum section lift coefficients greater than 2.0 were obtained and section lift-drag ratio at a lift coefficient of 1.0 (climb condition) varied from about 65 to 85 as the Reynolds number increased from about 2.0 million to 6.0 million.

Fluctuating Lift and Moment Coefficients for Cascaded Airfoils in a Nonuniform Compressible Flow

Abstract: The effects of compressibility on both the fluctuating lift and the fluctuating moment coefficients for cascaded airfoils due to an upstream nonuniformity are determined by obtaining a solution to the time-dependent, compressible, two-dimensional partial differential equation which describes the perturbation velocity potential. This is accomplished through an application of Fourier-trans form theory, with the resulting integral solution equation evaluated numerically by a matrix-inversion technique. The results presented show the variation in both the fluctuating lift and the fluctuating moment coefficients over the mean cascade inlet Mach number range of 0.0 (incompressible) to 0.9 with the cascade solidity, cascade stagger angle, interblade phase angle and reduced frequency as parameters.

Surface dipole strength by cross‐correlation method

Abstract: A new correlation technique enables the determination of local acoustic source strength on the surface of flat plates and simple aerofoil shapes. The method is based on Curle's equation for surface‐generated noise and involves cross‐correlation between local surface pressure and the farfield acoustic pressure. Flat plate airfoils of circular planform were positioned in a “quiet” open jet airflow. The distribution of dipole strength was obtained by moving the point of surface‐pressure measurement around on the surface. Distinctively different distributions were obtained for cases of incident turbulence (airfoil in turbulent mixing layer), separated flow (airfoil at 16° angle of attack in core of jet), and vortex shedding (airfoil at 0° angle of attack in core of the jet). The popular notion that noise comes predominantly from the edges of the plate was not strongly supported. In all cases, however, definite localized source regions were identified. The technique also enabled estimates of the correlation ar...

Unsteady Thin-Airfoil Theory for Subsonic Flow

Abstract: An approximate theory has been developed for the unsteady motion of a two-dimensional thin airfoil in subsonic flow. The basic theory is restricted by the condition that (s//?2)2 <^ 1, where e is i/(2n) times the product of the Mach number M and the conventional reduced frequency CM, and ft2 = 1 —M 2. Closed-form expressions are presented for the forces on, the circulation about, and the strength of the vortex wake emanating from a two-dimensional thin airfoil that is subjected to the general class of oscillating upwash distribution whose A dependence may be expanded in a cosine series. Expressions are also set down for three particularly important examples of this upwash distribution, namely the Kemp-type upwash, the convected sinusoidal gust, and the flutter case. Some preliminary comparisons of the present method with the few existing theories for the sinusoidal-gust case reveals good agreement up to at least M = 0.6 and for relatively large values of e.

Airfoil camber change system

Abstract: An airfoil camber change system for changing the camber of the leading or trailing edges of airfoils. The system includes a flexible continuous upper surface, a lower surface comprising a plurality of slidable overlapping segments and one or more actuation mechanisms. The actuation mechanism includes a plurality of bell cranks and links that are operatively connected to the upper and lower airfoil surfaces. A primary actuator is provided to drive the actuation mechanisms. When the system is driven the actuation mechanism changes the camber profile and maintains proper separation and support of the upper and lower surfaces. The profile is changed by bending the upper constant length surface and shortening the lower surface by increasing the overlap of the slidable overlapping segments.

Inviscid Flowfield of an Unsteady Airfoil

Abstract: Simple formulas have been developed from thin airfoil theory to describe the detailed inviscid, incompressible flowfield of an unsteady airfoil with thickness and camber. The solutions allow the various physical aspects of the problem and the parameters of the unsteady motion to be identified easily. The results agree well with numerical calculations and pressure measurements. The unsteady phase lag and attenuation of the inviscid pressure gradients near the leading edge explain the dynamic delay in laminar boundary-layer separation on oscillating airfoils, but not the characteristics of dynamic stall.

Exact Method of Designing Airfoils With Given Velocity Distribution in Incompressible Flow

Abstract: The inverse problem of airfoil theory, i.e., from a given surface velocity distribution determine the airfoil shape, is solved by conformal mapping procedures. The method is based upon prior work by Arlinger, which in turn is an extension of Lighthill's basic development. It involves the use of least squares and Lagrangian multipliers to modify the prescribed velocity distribution along a portion of the lower surface of the airfoil, thus ensuring that the modifications required for profile closure are minimized. The method developed should be of particular importance for calculating the shapes of new types of airfoils with high design lift coefficients, i.e., under conditions when conventional linearized theory breaks down. The method is exact in the sense of potential flow theory. Sample calculations are presented for a prescribed velocity distribution having an upper-surface constant-velocity region, followed by a Stratford-type zero-skin-friction portion, designed for Reynolds number = 3.10 and turbulent flow on both upper and lower surfaces.

Composite fan blade and method of construction

Abstract: A composite fan blade is fabricated by bonding two complementary preformed outer shell halves defining an airfoil planform of composite material sandwiching a metal spar which spar extends beyond the airfoil planform to form the root of the blade.

Airfoil with extendible and retractable leading edge

Abstract: A leading edge device (referred to as a foreflap) is formed by passing a curved parting line through an airfoil in such a manner as to allow the front part of the airfoil to rotate about a pivot which is close to the wing external contour. When the foreflap is thus extended, the resulting airfoil contour has a large bulbous nose which is formed mainly by the parting line. The exact contour of the resulting nose may be shaped to optimum aerodynamic contour for high lift operation without adversely affecting high speed performance, since the parting surface is inside of the clean airfoil when the foreflap is retracted. The foreflap may be operated by any conventional means, e.g., a simple hinge located at the pivot outside of the airfoil contour.

Stall-Flutter Analysis

Abstract: An analysis of unsteady airfoil stall and stall flutter is presented that is based on a series of approximations. Unsteady aerodynamic characteristics are related theoretically to static aerodynamic characteristics. Preliminary results show good agreement with experimental dynamic stall data. The analysis is applied to determine the boundaries for stall flutter, particularly for the straight wing of one candidate space-shuttle configuration. As formulated, the analysis should provide a conservative estimate - i.e., the predicted stall flutter region is slightly larger than the expected one, as demonstrated by comparison with experiments.

Non-hazardous ring airfoil projectile of non-lethal material

Abstract: A rotatable airfoil projectile comprising a hollow closed circular ring wing surrounding a central open area with a non-lethal riot control agent positioned within the hollow ring. The projectile consists of an aerodynamic lifting body of a thick ring wing geometry which uses spin imparted to it from a launching means for its gyroscopic stability. The combination of aerodynamic stability characteristics and high spin rate (i.e. above 2,000 rmp) results in a flat trajectory and extended range capability. The projectile ruptures on impact due to centrifugal and impact forces to distribute the non-lethal riot control payload about the target area. The sub-sonic launch velocity avoids bodily harm due to impact with a person even at point-blank range.

Variable camber multi-slotted flaps

Abstract: An airfoil is specified with a double slotted leading edge flap having two members and adapted to be used together with a multislotted trailing edge flap. The leading edge flap uses two slots for landing and one or no slots for take off. The rear leading edge member defines a slot substantially downstream of the cruise leading edge. Ahead of the trailing edge flap slot, there is provided a special slot across the wing''s surface in special cooperation with a spoiler. The airfoil surface between the downstream slot of the leading edge flap and the special slot near the trailing edge flap is short and with negligible camber. Special consideration is given to wing torsion, fuel volume, number of slots, and effect of varying airfoil thickness on flap geometric and aerodynamic parameters.

Reynolds Number Effects on the Shock Wave-Turbulent Boundary Layer Interaction at Transonic Speeds.

Abstract: : A Ludwieg tube experiment is described in which the pertinent features of the shock wave-boundary layer interaction on an airfoil are simulated with a two-dimensional flat plate in a supersonic nozzle. The nozzle is modified to impress an airfoil pressure distribution on the flat plate that is typical of a cruising flight condition. A normal shock wave is positioned at a fixed location on the plate, and measurements are made in the vicinity of the shock wave-boundary layer interaction zone.

Control surface deployment mechanism

Abstract: A deployment mechanism employing multiple, nonproportional, four-bar linkages for guiding the movement of an airfoil control surface. The linkages are connected at one of their ends to the airfoil control surface and at their other end to an airfoil. The connection points of each linkage to the airfoil define a set of lines which are skewed at progressively greater angles with respect to a reference line. The mechanism when activated guides the extension of the control surface into an overlapping position with the airfoil. The overlap forms a nozzle between the airfoil control surface and the airfoil surface. The amount of overlap at any point is a constant percentage of the airfoil chord length at that same point.

A Finite-Difference Method for Transonic Airfoil Design

Abstract: This paper describes an inverse method for designing transonic airfoil sections or for modifying existing profiles. Mixed finite-difference procedures are applied to the equations of transonic small disturbance theory to determine the airfoil shape corresponding to a given surface pressure distribution. The equations are solved for the velocity components in the physical domain and flows with embedded shock waves can be calculated. To facilitate airfoil design, the method allows alternating between inverse and direct calculations to obtain a profile shape that satisfies given geometric constraints. Examples are shown of the application of the technique to improve the performance of several lifting airfoil sections. The extension of the method to three dimensions for designing supercritical wings is also indicated.

Calculation of separated flows at subsonic and transonic speeds

Abstract: A boundary-layer integral approach is combined with a finite-difference relaxation method to calculate viscous interactions between separated flows at subsonic and transonic velocities Results are obtained for separated laminar flows on circular-arc airfoils at zero angle of attack and are compared with data of Collins (1972) Inviscid and viscous flows are covered

The effects of leading-edge serrations on reducing flow unsteadiness about airfoils.

Abstract: High frequency surface pressure measurements were obtained from wind-tunnel tests over the Reynolds number range 1.2 x 1,000,000 to 6.2 x 1,000,000 on a rectangular wing of NACA 63-009 airfoil section. A wide selection of leading-edge serrations were also added to the basic airfoil. Under a two-dimensional laminar bubble very close to the leading edge of the basic airfoil there is a large peak in rms pressure, which is interpreted as an oscillation in size and position of the bubble. The serrations divide the bubble into segments and reduce the peak rms pressures. A low Reynolds number flow visualization test on a hydrofoil in water was also conducted. A von Karman vortex street was found trailing from the rear of the foil. Its frequency is at a much lower Strouhal number than in the high Reynolds number experiment, and is related mathematically to the airfoil trailing-edge and boundary-layer thicknesses.

Effects of Delta Wing Separation on Shuttle Dynamics

Abstract: The unsteady aerodynamics of a candidate delta-wing shuttle orbiter have been investigated. Three potentially dangerous flow separation phenomena involving the delta wing have been identified, each of which could seriously compromise the flight dynamics. They are 1) leeside shock-induced separation, 2) sudden leading-edge stall, and 3) the subsonic leading-edge vortex and vortex burst. Each of these unsteady flow phenomena can be triggered by control surface deflection. Furthermore, wing stall and control-induced separation effects interact with the relatively large fuselage, increasing the coupling between lateral and directional stability characteristics. Trajectory shaping may be the most powerful means of dealing with these flow separation effects. The re-entry trajectory can be tailored to avoid or quickly traverse the unstable flow regions. However, it is prudent to use available means to control the duration and extent of separated flow when traversing the critical flow region. The unstable flow boundaries may be altered by modification of the wing planform, airfoil section, control deflections, etc.

Lift and Measurements in an Aerofoil in Unsteady Flow

Abstract: A wind tunnel is described which is capable of producing both “transverse” and “streamwise” gusts. An account is given of the lift and pressure fluctuations measured on an isolated aerofoil tested in the tunnel. The response to a transverse gust compares well with Kemp’s (1) theory although the pressure distribution is not as predicted. The results suggest that the wake behavior and in particular the existence of a separation region can in practice seriously affect the validity of applying the now classical unsteady vortex theory.Copyright © 1973 by ASME

Development of a nonlinear unsteady transonic flow theory

Abstract: A nonlinear, unsteady, small-disturbance theory capable of predicting inviscid transonic flows about aerodynamic configurations undergoing both rigid body and elastic oscillations was developed. The theory is based on the concept of dividing the flow into steady and unsteady components and then solving, by method of local linearization, the coupled differential equation for unsteady surface pressure distribution. The equations, valid at all frequencies, were derived for two-dimensional flows, numerical results, were obtained for two classses of airfoils and two types of oscillatory motions.

Study of the far wake vortex field generated by a rectangular airfoil in a water tank.

Abstract: Underwater towing experiments were carried out with a rectangular airfoil of aspect ratio 5.3 at 4 and 8 deg angles of attack and at chord-based Reynolds numbers between 2 x 100,000 and 7.5 x 100,000. Quantitative measurements by means of the hydrogen bubble technique indicated lower peak swirl velocities in the range of 100 to 1000 lenghts downstream than have been measured in wind tunnel of flight tests. The maximum circumferential velocity decayed whereas the turbulent eddy viscosity increased. This behavior and other known rates of vortex decay are explained in terms of an analytical solution for the vortex problem with time varying eddy viscosity. It is shown that this case corresponds to nonequilibrium turbulent vortex flow.

Two-Dimensional Subsonic Experiments with an NACA 0012 Airfoil

Abstract: : Two-dimensional tests were made with a 6-inch chord NACA 0012 section in the Calspan 8-foot Transonic Wind Tunnel to obtain section data with minimum wall interference effects. The measurements included three - component force data, surface pressure distributions and oil flow observations, and were made at Mach numbers ranging from 0.4 to 0.95 and at a chord Reynolds number of 1,000,000. Comparisons of the lift curve data with available theory indicate that the results are within about 1% of being two-dimensional. (Modified author abstract)

Blade for fluid flow machines

Abstract: The disclosure of this invention relates to a rotor blade for a fan for a gas turbine engine. The blade has in succession an aerofoil, a shank and a root, and the shape of the aerofoil is continued into the shank but at progressively reducing camber so that, whereas adjacent the aerofoil the camber of the shank is the same as that of the aero-foil, adjacent the root the camber of the shank is zero. This leads to a reduction in the stresses produced by a torsion couple arising in operation and acting in the sense tending to reduce the twist of the blade.