Browsing by Subject "aerodynamics"
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Item An Efficient Nonlinear Structural Dynamics Solver for Use in Computational Aeroelastic Analysis(2011-08-08) Freno, Brian AndrewAerospace structures with large aspect ratio, such as airplane wings, rotorcraft blades, wind turbine blades, and jet engine fan and compressor blades, are particularly susceptible to aeroelastic phenomena. Finite element analysis provides an effective and generalized method to model these structures; however, it is computationally expensive. Fortunately, these structures have a length appreciably larger than the largest cross-sectional diameter. This characteristic is exploitable as these potential aeroelastically unstable structures can be modeled as cantilevered beams, drastically reducing computational time. In this thesis, the nonlinear equations of motion are derived for an inextensional, non-uniform cantilevered beam with a straight elastic axis. Along the elastic axis, the cross-sectional center of mass can be o set in both dimensions, and the principal bending and centroidal axes can each be rotated uniquely. The Galerkin method is used, permitting arbitrary and abrupt variations along the length that require no knowledge of the spatial derivatives of the beam properties. Additionally, these equations consistently retain all third-order nonlinearities that account for flexural-flexural-torsional coupling and extend the validity of the equations for large deformations. Furthermore, linearly independent shape functions are substituted into these equations, providing an efficient method to determine the natural frequencies and mode shapes of the beam and to solve for time-varying deformation. This method is validated using finite element analysis and is extended to swept wings. The importance of retaining cubic terms, in addition to quadratic terms, for nonlinear analysis is demonstrated for several examples. Ultimately, these equations are coupled with a fluid dynamics solver to provide a structurally efficient aeroelastic program.Item An Examination of Configurations for Using Infrared to Measure Boundary Layer Transition(2012-10-19) Freels, Justin ReedInfrared transition location estimates can be fast and useful measurements in wind tunnel and flight tests. Because turbulent boundary layers have a much higher rate of convective heat transfer than laminar boundary layers, a difference in surface temperature can be observed between turbulent and laminar regions of an airfoil at a different temperature than the free stream air temperature. Various implementations of this technique are examined in a wind tunnel. These include using a heat lamp as an external source and circulating fluid inside of the airfoil. Furthermore, ABS plastic and aluminum airfoils are tested with and without coatings such as black paint and surface wraps. The results show that thermal conduction within the model and surface reflections are the driving issues in designing an IR system for detecting transition. Aluminum has a high thermal diffusivity so is a poor choice for this method. However, its performance can be improved using an insulating layer. Internal fluid circulation was far more successful than the heat lamp because it eliminates the reflected IR due to the heat lamp. However, using smooth surface wraps can mitigate reflection issues caused by the heat lamps by reducing the scatter within the reflection, producing an IR image with fewer contaminating reflections.Item Boundary-Layer Stability and Transition on a Flared Cone in a Mach 6 Quiet Wind Tunnel(2013-05-15) Hofferth, Jerrod WilliamA key remaining challenge in the design of hypersonic vehicles is the incomplete understanding of the process of boundary-layer transition. Turbulent heating rates are substantially higher than those for a laminar boundary layer, and large uncertainties in transition prediction therefore demand conservative, inefficient designs for thermal protection systems. It is only through close collaboration between theory, experiment, and computation that the state of the art can be advanced, but experiments relevant to flight require ground-test facilities with very low disturbance levels. To enable this work, a unique Mach 6 low-disturbance wind tunnel, previously of NASA Langley Research Center, is established within a new pressure-vacuum blow-down infrastructure at Texas A&M. A 40-second run time at constant conditions enables detailed measurements for comparison with computation. The freestream environment is extensively characterized, with a large region of low-disturbance flow found to be reliably present for unit Reynolds numbers Re < 11?10^6 m-1. Experiments are performed on a 5? half-angle flared cone model at Re = 10?10^6 m-1 and zero angle of attack. For the study of the second-mode instability, well-resolved boundary-layer profiles of mean and fluctuating mass flux are acquired at several axial locations using hot-wire probes with a bandwidth of 330 kHz. The second mode instability is observed to undergo significant growth between 250 and 310 kHz. Mode shapes of the disturbance agree well with those predicted from linear parabolized stability equation (LPSE) computations. A 17% (40 kHz) disagreement is observed in the frequency for most-amplified growth between experiment and LPSE. Possible sources of the disagreement are discussed, and the effect of small misalignments of the model is quantified experimentally. A focused schlieren deflectometer with high bandwidth (1 MHz) and high signal-to-noise ratio is employed to complement the hot-wire work. The second-mode fundamental at 250 kHz is observed, as well as additional harmonic content not discernible in the hot-wire measurements at two and three times the fundamental. A bispectral analysis shows that after sufficient amplification of the second mode, several nonlinear mechanisms become significant, including ones involving the third harmonic, which have not hitherto been reported in the literature.Item Flow control optimization in a jet engine serpentine inlet duct(2009-05-15) Kumar, AbhinavComputational investigations were carried out on an advanced serpentine jet engine inlet duct to understand the development and propagation of secondary flow structures. Computational analysis which went in tandem with experimental investigation was required to aid secondary flow control required for enhanced pressure recovery and decreased distortion at the engine face. In the wake of earlier attempts with modular fluidic actuators used for this study, efforts were directed towards optimizing the actuator configurations. Backed by both computational and experimental resources, many variations in the interaction of fluidic actuators with the mainstream flow were attempted in the hope of best controlling secondary flow formation. Over the length of the studies, better understanding of the flow physics governing flow control for 3D curved ducts was developed. Blowing tangentially, to the wall at the bends of the S-duct, proved extremely effective in enforcing active flow control. At practical jet momentum coefficients, significant improvements characterized by an improved pressure recove ry of 37% and a decrease in distortion close to 90% were seen.Item Structural and Aerodynamic Interaction Computational Tool for Highly Reconfigurable Wings(2011-10-21) Eisenbeis, Brian JosephMorphing air vehicles enable more efficient and capable multi-role aircraft by adapting their shape to reach an ideal configuration in an ever-changing environment. Morphing capability is envisioned to have a profound impact on the future of the aerospace industry, and a reconfigurable wing is a significant element of a morphing aircraft. This thesis develops two tools for analyzing wing configurations with multiple geometric degrees-of-freedom: the structural tool and the aerodynamic and structural interaction tool. Linear Space Frame Finite Element Analysis with Euler-Bernoulli beam theory is used to develop the structural analysis morphing tool for modeling a given wing structure with variable geometric parameters including wing span, aspect ratio, sweep angle, dihedral angle, chord length, thickness, incidence angle, and twist angle. The structural tool is validated with linear Euler-Bernoulli beam models using a commercial finite element software program, and the tool is shown to match within 1% compared to all test cases. The verification of the structural tool uses linear and nonlinear Timoshenko beam models, 3D brick element wing models at various sweep angles, and a complex wing structural model of an existing aircraft. The beam model verification demonstrated the tool matches the Timoshenko models within 3%, but the comparisons to complex wing models show the limitations of modeling a wing structure using beam elements. The aerodynamic and structural interaction tool is developed to integrate a constant strength source doublet panel method aerodynamic tool, developed externally to this work, with the structural tool. The load results provided by the aerodynamic tool are used as inputs to the structural tool, giving a quasi-static aeroelastically deflected wing shape. An iterative version of the interaction tool uses the deflected wing shape results from the structural tool as new inputs for the aerodynamic tool in order to investigate the geometric convergence of an aeroelastically deflected wing shape. The findings presented in this thesis show that geometric convergence of the deflected wing shape is not attained using the chosen iterative method, but other potential methods are proposed for future work. The tools presented in the thesis are capable of modeling a wide range of wing configurations, and they may ultimately be utilized by Machine Learning algorithms to learn the ideal wing configuration for given flight conditions and develop control laws for a flyable morphing air vehicle.