The object of this work was to design and build equipment to measure the aerodynamic forces which will cause self-induced oscillations of a body having an unsymmetrical cross-section, such as a D-section. This self-induced oscillation has been called "stall flutter" or in electrical transmission lines, "galloping-transmission lines." It is hoped that this equipment will be used to gain basic Information about these aerodynamic forces.

No description available.

A method for predicting the unsteady, subsonic, aeroservoelastic response of a wing has been developed. The air, wing, and control surface are considered to be a single dynamical system. All equations are solved simultaneously in the time domain by a predictor-corrector method. The scheme allows nonlinear aerodynamic and structural models to be used and subcritical, critical, and supercritical aeroelastic behavior may be modeled without restrictions to small disturbances or periodic motions. A vortex-lattice method is used to model the aerodynamics. This method accounts for nonlinear effects associated with high angles of attack, unsteady behavior, and deformations of the wing. The vortex-lattice method is valid as long as separation or vortex bursting does not occur. Two structural models have been employed: a linear model and a nonlinear model which accounts for large curvature. Both models consider the flexural-torsional motion of an inextensional wing. / Master of Science

The design of modern high-performance aircraft is toward increased aerodynamic efficiency, decreased structural weight, and higher flight speeds. Preliminary designs often exhibit a flutter instability within the desired operating envelope of the aircraft. Passive methods which have been used to solve the flutter problem include added structural stiffness, mass balancing, and speed restrictions. These methods may result in significant weight penalties. Studies by Boeing (ref. 1) show that weight penalties as high as 2 to 4% of the total structural weight may be required to solve the flutter problem passively by increasing the structural stiffness. Therefore, there is considerable interest in alternative methods of increasing the flutter speed beyond the original unaided value. / Master of Science

Panel flutter is an important design consideration for vehicles traveling at supersonic speeds. Most theoretical analyses of panel flutter consider the motion of the panel to be described adequately by classical thin plate theory. In such a theory, transverse shear deformations are neglected which is a reasonable assumption for solid plates. For a sandwich panel, neglect of transverse shear deformations may not be a good assumption in flutter analysis inasmuch as studies have indicated that the vibration and buckling behavior of such panels can be affected significantly by shear deformations. An analysis which considers transverse shear deformations is presented in order to determine the effect of finite transverse shear stiffness on the flutter behavior of sandwich plates. The sandwich theory used is due to Libove and Batdorf. The essential feature of this theory is that straight line elements perpendicular to the undeformed middle surface remain straight and of the same length but are not necessarily perpendicular to the deformed middle surface. The aerodynamic loading on the panel is given by two-dimensional static aerodynamics. The adequacy of such an approximation has been demonstrated for panels rigid in shear and the mathematical simplicity allows closed-form solutions to be found. The analysis proceeds from consideration of the equilibrium of an infinitesimal element. If equations are written in terms of the deflection and two shear deformations for equilibrium of forces in the z direction and equilibrium. of moments about the x and y axis, three differential equations involving the three unknown displacements are obtained. This system of equations is of sixth order with constant coefficients, but for simple support boundary conditions on the streamwise edges an exact solution can be obtained. The associated characteristics equation can be factored into a fourth degree equation and a second degree equation; thus an analytical expression can be obtained for the characteristic roots. The solution just described is a general solution for the motion of a sandwich panel simply supported along streamwise edges and subject to inertia loading and aerodynamic forces given by two-dimensional static aerodynamics. Any combination of boundary conditions consistent with the sandwich plate theory used can be applied at the leading and trailing edges. Two cases are considered: simply supported leading and trailing edges and clamped leading and trailing edges. With the use of either set of boundary condition, a transcendental equation is obtained which is satisfied by various combinations of frequency and dynamic pressure. The dynamic pressure necessary to cause the frequency to become complex corresponds to divergent oscillatory motion or flutter. Values of the flutter dynamic pressure have been calculated as a function of length-width ratio for a large range of shear stiffness. For infinite shear stiffness the results agree with those established by previous investigators. As shear stiffness decreases, the flutter dynamic pressure usually decreases also. An unusual result of the analysis is that at low length-width ratios, a clamped panel has a lower flutter dynamic pressure than a simply supported panel even though the vibration frequencies are higher for the clamped panel. Results are not presented for panels with normal inplane loadings but they can be obtained from the equations given. The analysis shows that flutter is independent of normal inplane loadings perpendicular to the flow direction just as was found for panels rigid in shear. An approximate two-mode Galerkin solution to the problem has been obtained by a previous investigator. Comparison of the exact solution to the approximate solution shows the approximate analysis to be in increasing error as length-width ratio increases or shear stiffness decreases. / Ph. D.

This dissertation deals with the instability, known as flutter, of the lifting and control surfaces of aircraft of advanced design at high altitudes and speeds. A simple model is used to represent the aerodynamics for flutter analysis of a two-degree-of-freedom airfoil system. Flutter solutions of this airfoil system are shown to be algebraically homomorphic in that solutions about different elastic axes can be found by mapping them to those about the mid-chord. Algebraic expressions for the flutter speed and frequency are thus obtained. For the prediction of flutter of a wing at transonic speeds, an accurate and efficient computer code is developed. The unique features of this code are the capability of accepting a steady mean flow regardless of its origin, a time dependent perturbation boundary condition for describing wing deformations on the mean surface, and a locally applied three-dimensional far-field boundary condition for minimizing wave reflections from numerical boundaries. Results for various test cases obtained using this code show good agreement with the experiments and other theories.

No description available.

Thesis (M.S. in Aeronautical Engineering)--Naval Postgraduate School, September 2002. / Thesis advisor(s): Kevin D. Jones, Max F. Platzer. Includes bibliographical references (p. 61). Also available online.

Thesis (Ph. D.)--University of Alabama at Birmingham, 2006. / Description based on contents viewed Jan. 26, 2007; title from title screen. Includes bibliographical references (p. 112-126).

The word hysteresis is derived from a Greek word meaning "to lag 'behind". As specifically applied to fluid flow around bodies with transient angles of attack in and near the stall region, "aerodynamic hysteresis" is used to describe the effect of delay in boundary layer separation and reattachment upon the lift, drag and pitching moment. Experimental work done on airfoils by H. Studer showed that for increasing angles of attack flow "separation is delayed to an angle of attack appreciably greater than that for a stationary airfoil. On the return movement, re-establishment of a smooth flow is also delayed." [1]* The result is that under transient conditions "more than one value of flow coefficient (and thus lift, drag and pitching moment) can be obtained for a single angle of attack. . ., depending upon the direction in which the particular angle of attack is approached." [2]