No description available.

No description available.

The performance and stalling behavior of an axial-flow compressor subjected to several different inlet distortion patterns was investigated. The effect of inlet distortion on overall compressor performance was determined through the measurement of compressor characteristics for each inlet flow condition. Dynamic pressure transducers were employed to investigate rotating stall cell behavior during the inception of stall. Rotor blade response to distorted inflow was measured in the form of average blade pressure profiles by using a scanning valve. Results indicated a substantial reduction in total pressure rise capability for distorted operation. A 25 per cent loss in stall pressure rise was observed for the most severe distortion level. The stall cell was found to rotate in the direction of rotor motion, but at one-half the rotor speed. The cell encompassed the rotor blade tip region down to approximately midspan. During the onset of stall, the circumferential extent of the cell was observed to vary from 60 to 80 degrees. At the rotor blade tip the stall cell relative pressure fluctuations indicated zero flow through the cell. The amplitude of the stall cell was attenuated in the distorted flow region due to the lower air velocity behind the distortion screens. Rotor blade suction side pressure measurements indicated that increasing the circumferential extent of distortion above some "critical" value induced blade stall at higher flow rates. For the low speed compressor used, it appears that the critical angle phenomena is a function of compressor design and is independent of distortion level. / Master of Science

Extensive hot-wire measurements were conducted in an incompressible turbulent flow around a wing-body junction. The measurements were performed adjacent to the body and up to 11.56 chord lengths downstream of the body. The junction wake flow entered an adverse pressure gradient region approximately 6 chord lengths downstream. This region's geometry approximated the aft portion of an aircraft fuselage or a submersible's hull. The body geometry was formed by joining a 3:2 elliptic nose to a NACA 0020 tail section at their respective maximum thickness locations. The author's measurements were taken with approach flow conditions of Reθ = 6,300, and δ/T = .513, where T is the maximum body thickness. The results clearly show the characteristic horseshoe vortex flow structure. The vortex flow structure is elliptically shaped, with â (W)/â Y forming the primary component of streamwise vorticity. Near wall measurements show a thin layer of highly concentrated vorticity, underneath and opposite in sign to the primary vortex, which is created by the wall no-slip condition. The development of the flow distortions and associated vorticity distributions are highly dependent on the geometry-induced pressure gradients and resulting flow skewing directions. A quantity known as the "distortion function" was used to separate the distortive effects of the secondary flow from those of the body and the local "2-D" boundary layer. The distortion function revealed that the adverse pressure gradient flow distortions grew primarily because of the increasing boundary layer thickness. The author's results were compared to several other data sets obtained using the same body shape, enabling the determination of the approach boundary layer effects. The primary secondary flow structure was found to scale on T in the vertical and cross-stream directions, revealing that the juncture flow is driven by the appendage geometry and associated pressure gradients. A parameter known as the momentum deficit factor (MDF = (ReÏ ) 2 (θ/T) was found to correlate the observed trends in mean flow distortion magnitudes and vorticity distribution. Variations in flow skewing were observed to be comparable to changes in MDF, suggesting that this flow parameter changes the effective skewing magnitudes around a wing-body junction. Mean flow distortions were found to increase with decreasing values of MDF. A numerical study was also performed to gain additional insights into the effects of appendage nose geometry. The velocity distributions around approximately 30 different appendage cross-sections were estimated using 2-D potential flow calculations. A correlation was found between the appendage nose bluntness and the average vortex stretching rate, and also between the invisicid velocity distribution and an experimentally determined non-dimensional circulation estimate. / Master of Science

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No description available.

No description available.

No description available.

The paraglider wing consists of leading-edge booms and a keel boom joined together at the nose and a flexible sail whose surface carries the aerodynamic pressure loading. The payload is suspended beneath the wing by cables which are used to control the wing. Adjusting the length of these cables controls the glide path of the wing by shifting the position of the payload with respect to the wing. The deflected shape of the sail depends on the pressure distribution over the sail; the pressure distribution in turn depends aerodynamically on the deflected shape of the sail. It is the purpose of this thesis to derive the equilibrium equations for the sail and to integrate these equations to find expressions for the stress resultants in terms of the pressure on the sail and the deflected shape of the sail. By integration of these expressions for stress resultants, along the edges of the sail, the resultant forces applied by the sail to the leading-edge booms and keel boom are found. Then, by considering the streamwise components and the components normal to the stream of the boom forces, the drag and lift forces are obtained. These expressions for the lift and drag forces, and for the boom forces, are given in terms of the boundary value of the stress resultants and can be applied for any aerodynamic theory appropriate to the speed range being considered. When the appropriate aerodynamic relationship, between pressure and deflected shape, is substituted into the boundary, conditions for stress resultant at the trailing edge of the sai1, the criterion for determining the deflected shape is obtained. Once the deflected shape is known all the other quantities can be determined. In order to show an application of the analysis, the equations were specialized for Newtonian impact theory. This theory yields a simple pressure-deflected relationship. This aerodynamic theory, which has found applications for hypersonic speeds, has been employed since it shows the application of the method in a simple manner. In addition, it has been previously for a rigid idealization of a paraglider wing and thus provides a ready means for comparison. Hence numerical results can be used to test the accuracy of the rigid idealization. These results showed that the deflected shape of the flexible paraglider wing differed considerable from the conical shape of the rigid wing over the complete range of angle of attack. The differences in shape result in different pressure distributions over the surface of the wing and as a result the lift and drag coefficients, and especially the lift-to-drag ratio, for the flexible wing were significantly different from the values for the rigid wing. The boom forces and the distribution of stress resultants over the surface of the sail were obtained. The stress resultants along radial lines were found to be proportional to the distance from the nose of the wing. The calculated stress resultants and boom forces provide a basis for design of sails, booms, shroud lines, and spreader bars for a paraglider for hypersonic flight. Effects of dihedral angle (raising or lowering of the leading-edge booms relative to the keel) were also considered. The pressure distributions, the lift and drag coefficients, and the ratio to drag were found for several dihedral angles at a given angle of attack. / Ph. D.

The research contributes toward a fully integrated multidisciplinary wing design synthesis by development of an appropriate structural model. The goal is to bridge the gap between highly idealized structural beam / aerodynamic strip models and the very detailed finite element and computational fluid dynamics, FEM/CFD, techniques. The former provides insufficient accuracy for flutter analysis of modern low aspect ratio composite wings. The latter is too computationally intensive for use in the inner loop of a simultaneous multidisciplinary optimization problem. The derived model provides a useful preliminary design tool as well. / Ph. D.