Thesis (M.S.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 1996. / Includes bibliographical references (p. 97-80). / by D.J. Orr, Jr. / M.S.

Thesis (Ph. D.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2010. / Cataloged from PDF version of thesis. / Includes bibliographical references (p. 223-231). / When considering the formation control problem for large number of spacecraft, the advantages of implementing control approaches with a centralized coordination mechanism can be outpaced by the risks associated with having a primary vital control unit. Additionally, in formations with a large number of spacecraft, a centralized approach implies an inherent difficulty in gathering and broadcasting information from/to the overall system. Therefore, there is a need to explore efficient decentralized control approaches. In this thesis a new approach to spacecraft formation control is formulated by exploring and enhancing the recent results on the theory of convergence to geometric patterns and exploring the analysis of this approach using the tools of contracting theory. First, an extensive analysis of the cyclic pursuit dynamics leads to developing control laws useful for spacecraft formation flight which, as opposed to the most common approaches in the literature, do not track fixed relative trajectories and therefore, reduce the global coordination requirements. The proposed approach leads to local control laws that verify global emergent behaviors specified as convergence to a particular manifold. A generalized analysis of such control approach by using tools of partial contraction theory is performed, producing important convergence results. By applying and extending results from the theory of partially contracting systems, an approach to deriving sufficient conditions for convergence is formulated. Its use is demonstrated by analyzing several examples and obtaining global convergence results for nonlinear, time varying and more complex interconnected distributed controllers. Experimental results of the implementation of these algorithms were obtained using the SPHERES testbed on board the International Space Station, validating many of the important properties of this decentralized control approach. They are believed to be the first implementation of decentralized formation flight in space. To complement the results we also consider a short analysis of the advantages of decentralized versus centralized approach by comparing the optimal performance and the effects of complexity and robustness for different architectures and address the issues of implementing decentralized algorithms in a inherently coupled system like the Electromagnetic Formation Flight. / by Jaime Luís Ramírez Riberos. / Ph.D.

Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, February 2002. / Includes bibliographical references (p. 76-79). / by Jessica J. Márquez. / S.M.

Thesis (Ph. D.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2006. / Includes bibliographical references (p. 104-106). / Most modem blade designs in axial-flow compressors diffuse the flow efficiently over 20% to 80% of blade span and it is the endwall regions that set the limits in compressor performance. This thesis addresses the estimation, control and mitigation of three-dimensional separation near the hub corner in axial- flow compressors. A simple method to estimate the onset of hub comer separation in compressor blade passages has been developed. A parameter is defined to quantify the combined effect of adverse pressure gradient and secondary flow which are the two main mechanisms contributing to the formation of three-dimensional flow separation. There is a critical value of the parameter at which the onset of three-dimensional flow separation occurs. Data from existing research and production compressors show the generality of the separation criterion. The new parameter captures the alleviating effect of boundary layer skew on three-dimensional flow separation. Using this concept, a flow control scheme has been developed to mitigate hub comer separation by injecting spanwise momentum from the blade suction surface. A proof of concept flow control experiment demonstrates a reduction in stagnation pressure loss coefficient of 8% with an injection flow of 0.8% of the cascade mass flow. / by Vai-Man Lei. / Ph.D.

Thesis (M.S.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 1994. / Includes bibliographical references (p. 48-51). / by Kathryn A. Fricks. / M.S.

Thesis (M.S.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 1995. / Includes bibliographical references (leaves 197-204). / by David Shia. / M.S.

Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2011. / Cataloged from PDF version of thesis. / Includes bibliographical references (p. 263-267). / A stepwise methodology to assist in the development of a model to determine the interlaminar stress fields and the mechanisms that give rise to these stress fields in transversely-loaded grooved composite laminates was developed. The methodology consists of five steps, with each step representing a laminate configuration with an increase in complexity from the previous step. Models for each step were developed in the context of a general formulation proposed in the literature that was previously used to analyze the proposed problem for Step 1. This general formulation is based around an assumed stress approach, where unknown coefficents in the assumed stress shapes are solved via the Principle of Minimum Complementary Energy. Models derived from such an approach were designed to be both accurate in the results as well as more efficient in runtime as compared to other models (e.g. finite element models). Such models are particularly useful for preliminary design, where various laminate configurations need to be analyzed efficiently to find designs for more detailed analysis and further modification. The results from the models for Steps 1 through 3 were found to be in good agreement with results in the literature, when available, or finite element results when analyzing configurations with results not found in literature. Characteristic results were taken from the models for Steps 1 through 3 in order to determine modifications, as well as identifying phenomena in the stress distributions, that require particular attention in the formulation of subsequent models. Key controlling factors in the model for Steps 1 through 3 are identified from the results. Issues encountered in Step 4 with regard to representation of slanted dropoffs prevent further model development. The feasibility of the formulation of the Step 5 model, involving loaded edges, was established assuming a working Step 4 model. Modifications of Steps 4 and 5 of the initially-proposed stepwise methodology were developed such that models for the proposed problems can be developed using the general formulation established for previous steps. These modified steps allow for a closed-form solution within the context of the general formulation as well as identification of the mechanisms and laminate parameters that affect interlaminar stress fields in transversely-loaded grooved laminates. Recommendations for future work are made. / by Alberto Joseph Gayon. / S.M.

Thesis (Ph.D.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2000. / Includes bibliographical references (p. 177-186). / A novel new actuator for helicopter rotor control, the X-Frame Actuator, was developed, demonstrating superior performance for applications requiring compact, fast acting, large stroke actuation. The detailed experimental characterization of this actuator is described, including bench-top output energy measurements and transverse shake test performance. A Mach scaled rotor blade utilizing the X-Frame actuator to power a trailing edge servo-flap near the tip was also designed, manufactured and tested. A description of the design and composite manufacturing of the rotor blade and servo-flap is presented. Preliminary bench tests of the active blade actuation system are also presented. The hover tests of the active blade provided transfer function identification of the performance of the actuator in producing flap deflections, and the response of the rotor from deflections of the servo-flap. At the highest field level of 60 V/mil P-P the actuation system produces 7.75 degrees of quasi-static peak-to-peak flap deflection in hover. The servo-flap produces significant control authority, especially near the 3/rev frequency that would be important for the CH-47. Scaled to a full-sized CH-47, the rotor can produce over 16,000 lb peak-to-peak thrust variation at 3/rev, which is 32% of the aircraft's gross weight. Closed-loop feedback control was experimentally applied to the model rotor system. Both single frequency and combined frequency controllers were successfully implemented on the rotor. Most significantly, simultaneous control of 1/rev, 3/rev, 4/rev, 5/rev, and 6/rev harmonic vibration has been successfully demonstrated. The peak vibrations were eliminated at each frequency, as well as the vibrations over a small bandwidth surrounding each peak. Experimental comparison of continuous time versus discrete time control has shown the former to be a more effective approach for vibration reduction. / by Eric Blade Prechtl. / Ph.D.

Massachusetts Institute of Technology. Dept. of Aeronautics and Astronautics. Thesis. 1966. M.S. / Bibliography: leaf 42. / M.S.

Massachusetts Institute of Technology. Dept. of Aeronautics and Astronautics. Thesis. 1969. M.S. / MICROFICHE COPY ALSO AVAILABLE IN AERO LIBRARY. / Bibliography: leaf 61. / M.S.