Air flow in the wake region of a two-dimensional (plane) body with a blunt base has been studied using numerical simulations. The objective of this study is (1) to observe the behavior of large dynamic structures in the plane wake at several Mach numbers from low (almost incompressible) up to M = 2.46 and examine their effect on the base pressure, and (2) to address the nature of the instability in the shear layers bounding the wake flow at M = 2.46 and observe the structures that arise from this instability. A code was developed for this study which solves the compressible Navier-Stokes equations in two or three dimensions. This code may be used for either Direct Numerical Simulations (DNS) or Large Eddy Simulations (LES). A spatial model is used, with the computational domain arranged around the trailing edge of a two-dimensional flat plate with a blunt base. Two-dimensional simulations were carried out at Mach numbers of M = 0.25, M = 1.20, and M = 2.46. At all Mach numbers, the flow was found to be unstable with respect to sinuous (antisymmetric) disturbances, with the critical Reynolds number increasing with increasing Mach number. These disturbances grow to a periodic state, and a Karman vortex street is formed. Examination of the supersonic cases revealed that expansion fans in the flow at the corners are the primary cause of the low base pressure, and that disruptions in the expansions raise the base pressure. At M = 2.46 and Reynolds numbers starting at Re = 100, 000, an intermittent shear layer instability was also found, excited by sinuous disturbances. The two instability 2 modes interact to produce a chaotic behavior. Above Re = 200, 000, the shear layer instability appears close to the base without sinuous disturbances, forming rows of vortices in the shear layers. Preliminary three-dimensional simulations were carried out at M = 2.46, examining the variation in the growth rate of three-dimensional disturbances with spanwise wavelength.

Orbital debris is a growing concern for all space applications. Specifically, the ASPOD concept has been proposed to help eliminate the debris population. This dissertation focuses on the characterization and feasibility of processing (cutting) structural metals in low earth orbit using concentrated solar energy as they pertain to the ASPOD concept. In characterizing the process, both experimental and analytical techniques were utilized. Analytically, a non-linear explicit finite difference model was created that examined how the heat transfer and physical parameters affect metal processing in low earth orbit. In addition, the model was used to develop a thermal criterion under which processing of aluminum debris can be accomplished with concentrated solar radiation. The experimental investigation entailed designing and constructing two experimental apparatuses. The first experimental apparatus was utilized to demonstrate the control of radiative surface properties on simulated orbital debris and to allow for the conceptual testing of physical parameters. The second experimental apparatus was constructed to demonstrate the entire cutting process. The feasibility of cutting structural members in low earth orbit with concentrated solar is discussed and demonstrated. Finally, the effect of these results on the ASPOD concept and threat of orbital debris is addressed.

An ellipsoidal dome is a conformal optical element used to replace a hemispherical dome on a missile to enhance its performance by reducing its aerodynamic drag. Conformal optics are a general class of optical systems in which the optical elements are shaped to optimize something other than image quality, such as aerodynamics. An ellipsoidal dome has lower aerodynamic drag than a comparably sized hemispherical dome. On a missile, lower drag improves its aerodynamic performance by increasing its range and fuel efficiency but degrades the quality of the transmitted wavefront. In particular, an ellipsoidal dome introduces a varying aberration component that depends on the orientation of the aperture stop, which is pivoted about a fixed axis inside the dome. The transmitted ray bundle is incident only on a portion of the dome surface, and the included area lacks axial symmetry. To better understand the imaging characteristics of an ellipsoidal dome in this application, the first- and third-order optical properties of a constant thickness dome are investigated. Particular emphasis is placed on the geometry and symmetry of an ellipse, which impose certain constraints on the form of the aberration coefficients. The geometry is defined in terms of the aerodynamic fineness ratio, outer diameter, and center thickness of the dome. Emphasis is placed on third-order astigmatism and coma, which are shown to be the dominant aberration terms. The effects of varying the fineness ratio, thickness, and index of refraction of a dome are also investigated.

Research is presented on the development and testing of a new procedure for the time dependent, spatially varying numerical simulation of wall bounded turbulent flows. The Flow Simulation Methodology (FSM), as it is now known, was originally proposed by Speziale (1996a) for the purpose of computing complex, non-equilibrium flows which are currently beyond the reach of Smagorinsky based Large-Eddy Simulations (LES). The new method represents a hybrid approach that combines favorable aspects of Reynolds stress modeling [used for Reynolds Averaged Navier-Stokes (BANS) calculations] with the underlying principles of LES. For instance, Reynolds stress models developed for non-equilibrium, anisotropic, and/or rotational flows can be utilized in the unsteady manner of LES, i.e. where the flow field is decomposed into resolved-scale (calculated) and subgrid-scale (modeled) components, thereby reducing computational requirements. The key to the FSM is a contribution function which provides a degree of local turbulence modeling that is dependent upon the ratio of the numerical resolution to the Kolmogorov length-scale, an estimate for the smallest scales of turbulent motion. With this approach, a calculation resolved to the level of a Direct Numerical Simulation (DNS) can proceed continuously to a Reynolds Averaged Navier-Stokes calculation as the numerical resolution is decreased and/or the Reynolds number is increased. In between these two limits, an "untraditional" LES is recovered. The method is untraditional because it replaces the commonly employed Smagorinsky subgrid-scale model, which is known to have considerable limitations, with a more capable Reynolds stress model. A detailed evaluation of the Flow Simulation Methodology is made for the test case of a transitional and turbulent flat plate boundary layer with zero pressure gradient. The relatively simple geometry is chosen because the technical issues associated with combining elements of RANS calculations and LES must be established and the FSM itself must be validated before more complex flows can be attempted. The Reynolds stresses needed for the new method are computed using the two-equation Algebraic Stress Model (ASM) of Gatski & Speziale (1993) developed for non-equilibrium turbulent flows. Results of FSM calculations are compared with results obtained from coarse grid DNS, traditional LES based on the Smagorinsky subgrid-scale model, and RANS, all of which are implemented using an identical core computer code. This approach is extremely valuable to the evaluation of the FSM since a common code allows for certain behaviors to be more easily attributed to the turbulence models as opposed to numerical effects. Further validation is achieved through comparisons of FSM results with various direct numerical simulations and experiments available in the literature.

The acoustical signatures as observed by an auditor on the ground are explored for various radiating bodies. Specifically, a theory that describes the origin of sonic booms of two unsteady point sources and of an airplane is developed. In 1968, Garrick and Maglieri conducted an experiment where a General Dynamics F-106 was subjected to sinusoidal pitch oscillations. At the time, the results of the observed sonic boom were not understood; they had expected a distorted sonic boom. The theory presented in the present study offers an explanation of the results. An essential point needed in order to understand their observations is the source distribution for an acoustically radiating body. This source distribution occupies a region of space many times the length of the airplane. Therefore, any attempts to distort a sonic boom must deal with the grand scale of the source distribution.

This thesis provides a general framework for the manual teleoperation of kinematically redundant space-based manipulators. It is proposed to break down the task of controlling the motion of a redundant manipulator into a sequence of manageable sub-tasks of lower dimension by imposing constraints on the motion of intermediate bodies of the manipulator. This implies that the manipulator then becomes a non-redundant kinematic chain and the operator only controls a reduced number of degrees of freedom at any time. However, by appropriately changing the imposed constraints, the operator can use the full capability of the manipulator throughout the task. / Also, by not restricting the point of teleoperation to the end effector but effectively allowing direct control of intermediate bodies of the robot, it is possible to teleoperate a redundant robot of arbitrary kinematic architecture over its entire configuration space in a predictable and natural fashion. / It is rigourously proven that this approach will always work for any kinematically redundant serial manipulator regardless of its topology, geometry and of the number of its excess degrees-of-freedom. Furthermore, a methodology is provided for the selection of task and constraint coordinates to ensure the absence of algorithmic rank-deficiencies. / Two novel algorithms are provided for the symbolic determination of the rank-deficiency locus of rectangular Jacobian matrices: the Singular Vector Algorithm and the Recursive Sub-Determinant Algorithm. These algorithms are complementary to each other: the former being more computationally efficient and the latter more robust. / The application of the methodology to sample cases of varying complexity has demonstrated its power and limitations: It has been shown to be powerful enough to generate complete sets of task/constraint coordinate pairs for realistic examples such as the Space Station Remote Manipulator System and a simplified version of the Special Purpose Dexterous Manipulator.

The separation of the boundary layer at the trailing edge of sails is considered in the present work. This problem was identified as one of the major reasons for the disagreement between existing theories and experimental results. The separation of the flow on flexible surfaces is a complex phenomenon because of the interdependence between the shape of the sail, the pressure distribution and the separation position. The problem is simplified by considering the flow over a two-dimensional sail set at design incidence to isolate the effect of the separation of the flow. / A theoretical model predicting the pressure distribution around a rigid aerofoil with trailing edge separation has been developed. The model uses a point source in potential flow to simulate the effect of the separated region on the flow. The separation position is determined using a turbulent boundary layer development calculation based on an integral method. The source strength and its position are modified iteratively until a set of closure conditions is satisfied. / The model is first tested on circular-arc aerofoils and validated against experimental results of obtained on five different circular-arc aerofoils of camber ratio ranging from 10% to 27%. The theoretical results are also compared with other experimental results found in the literature. The agreement between the viscous theory and the experimental results is satisfactory up to camber ratios of approximately 23%. / The theoretical model is then extended to simulate the flexibility of a two-dimensional sail. The shape of a rigid aerofoil is modified iteratively using the pressure distribution obtained from the flow separation model and the sail equation until convergence. / The results of the viscous theory for two-dimensional sails are compared with a new set of experimental results obtained on four different two-dimensional sails of excess-length ratios ranging from 0.097 to 0.167. The predictions of the theory for a lower range of excess-length ratios are compared with experimental results from other authors. For the two-dimensional sails the viscous theory is generally in good agreement with experimental measurements up to excess-length ratios of 0.167. / The theory presented in the present thesis demonstrates that a single source can simulate the essential effect that separation of the boundary layer near the trailing edge has on a thin cambered aerofoil. The integration of the separation model in an iterative scheme made it possible to determine the effect of trailing edge separation on the aerodynamic performance of a two-dimensional sails at design incidence.

<p> A low-speed wind tunnel experiment was performed to study the effect of a backward-facing step on transition in a swept-wing flow. Detailed hot-wire measurements were used to assess the flow field characteristics on a swept flat plate with and without a backward-facing step. A pressure body was installed on the ceiling to induce a pressure field simulating that of an infinite swept wing. The step height was approximately 50% of the boundary-layer thickness at the step. Measurements without the step confirmed the dominance of the stationary crossflow instabilities leading to a high-frequency secondary-instability breakdown. The backward-facing step had a local destabilizing effect on the growth of the dominant stationary crossflow mode and the harmonic of the dominant mode. The stationary crossflow disturbances reached small amplitudes (3 to 5% U_e) before breakdown occurred. The transition front moved forward as the initial amplitude of the stationary crossflow disturbance was increased. The step introduced a flow field rich with unsteady disturbances. Three different families of unsteady disturbances were identified corresponding to three distinct frequency bands in the 80 to 1500 Hz range. Wave angles and phase speeds were measured for each type of disturbance. The disturbances are believed to correspond to a traveling crossflow-type disturbance, a TS-type disturbance, and a free shear layer instability. Each of the disturbances were modulated through interaction with the stationary crossflow modes. The spanwise modulation was different for each family and was seen in the distortion of the amplitude and phase. Larger stationary crossflow vortices resulted in larger peak amplitudes of the unsteady disturbances at similar streamwise locations. The mean-flow modulation appears to affect the local stability of the unsteady disturbances even at low stationary-crossflow amplitudes. The local destabilization of the unsteady disturbances is believed to be responsible for the sensitivity of transition location to stationary crossflow amplitude. Breakdown was initiated despite the low amplitude of the unsteady disturbances (2 to 4% U_e). Nonlinear interactions were observed between the different unsteady disturbances and may be ultimately responsible for breakdown to turbulence.</p>

When a detonation propagates in an explosive that partially fills a channel, a shock wave can be generated in the air gap between the explosive and its confinement. This is typically referred to as the channel effect. In the air gap, the detonation products behave as a piston, which drives the precursor air shock wave ahead of the detonation. Since the shock wave runs ahead of the detonation, it preconditions the explosive and the detonation propagates into shocked explosive. This can affect the detonation in various ways, depending on the nature of the explosive. Properties such as heterogeneity, porosity and sensitivity of the explosive will determine how the precursor shock will affect the detonation propagation. Four different coupling mechanisms have been identified and are discussed in this thesis. They have been called: precompression, detonation initiation, surface ignition and dead pressing. / In the present study, three cases are investigated experimentally: precompression, detonation initiation and the case where no coupling occurs. The goal is to elucidate the respective propagation mechanisms. It is found that boundary layers on the channel walls significantly affect the precursor shock wave propagation. This effect is modeled and the results are compared to experiments. / When the explosive is PETN powder, the detonation is found to accelerate to 1.5 times the Chapman-Jouguet velocity. Experiments performed indicate that this is due to precompression of the PETN. Again this effect is modeled and compared with experiments. / It is also demonstrated that in the present experiments, coupling via initiation does not occur. However, experiments were performed to determine why the explosive is not initiated by the precursor shock wave. It is found that the initiation delay for the strength of precursor shock generated is simply too long for any coupling to occur.

Modal testing is often employed in the determination of natural frequencies and damping levels in aircraft structures. In aircraft flutter testing, potentially dangerous flight regimes are avoided by obtaining modal frequency and damping values at airspeeds well below the flutter speed and extrapolating the data to estimate the airspeed at which the onset of flutter instabilities is expected to occur. In the modal analysis, the structure is typically assumed to be linear and the parameters to be time-invariant. Nonlinearities in aeroelastic systems can arise from both structural and aerodynamic sources and may initiate aeroelastic instabilities both above and below the flutter speed predicted by linear theory. Typical nonlinear responses include limit cycle oscillations and in some cases, chaotic response. For aeroelastic systems containing even small nonlinearities, the nonlinear frequency response curve may be distorted, and this distortion can contribute to errors in the values of frequency and damping obtained during modal testing. The current study includes an analytical and an experimental investigation into the modal testing of a nonlinear aeroelastic system. / In the case of aeroelastic systems containing limited structural nonlinearities, the nonlinearity, although it changes the system frequency and damping values and distorts the transfer function, does not substantially affect the critical flutter speed. For this reason, the nonlinearity behaves a little like "noise" in that it prevents accurate values of frequency and damping from being obtained during the modal test. One solution to this problem is to separate the linear and nonlinear portions of the frequency response using spectral decomposition methods. In the analytical portion of this study, a specific spectral decomposition technique is tested on numerical data, and the results show that the technique may be used to separate the linear and nonlinear portions of the transfer function obtained from the nonlinear aeroelastic system response to a random forcing input. / In the experimental portion of the study, subsonic wind tunnel experiments are performed on a two degree-of-freedom wing section with a freeplay-type nonlinearity in the pitching degree-of-freedom. The experiments demonstrate the effect of the freeplay on the aeroelastic response, including the presence of limit cycle flutter for specific parameter combinations. The effects of variations in both freeplay length and frequency ratio of the underlying linear system are examined for both the damped and the limit cycle response. Time histories of the damped response are used to estimate frequency and damping values, and to predict critical flutter speeds. The amplitude and frequency of the LCO response is presented for three different freeplay lengths and five frequency ratios. / The experimental setup is modeled analytically, and the equations solved numerically to obtain time-history data for the free response of the airfoil to initial displacements in each of its two degrees-of-freedom. The results are used to validate the mathematical model of the aeroelastic system subject to subsonic flow. Numerical results are obtained for the transient behaviour of the system including the damped and limit cycle responses observed during the experimental portion of the study. The model is used to demonstrate the influence of frictional forces on the response of the experimental system. / In conclusion, three additional techniques of nonlinear dynamical analysis are investigated for their potential within the context of the nonlinear aeroelastic modal analysis problem. Two of the methods are tested on the experimental time-history data, while the third is shown to be applicable to the simulated data for the forced response of the system.