Thesis (M.S.) -- Mississippi State University. Department of Aerospace Engineering. / Title from title screen. Includes bibliographical references.

abstract: During the past five decades neurosurgery has made great progress, with marked improvements in patient outcomes. These noticeable improvements of morbidity and mortality can be attributed to the advances in innovative technologies used in neurosurgery. Cutting-edge technologies are essential in most neurosurgical procedures, and there is no doubt that neurosurgery has become heavily technology dependent. With the introduction of any new modalities, surgeons must adapt, train, and become thoroughly familiar with the capabilities and the extent of application of these new innovations. Within the past decade, endoscopy has become more widely used in neurosurgery, and this newly adopted technology is being recognized as the new minimally invasive future of neurosurgery. The use of endoscopy has allowed neurosurgeons to overcome common challenges, such as limited illumination and visualization in a very narrow surgical corridor; however, it introduces other challenges, such as instrument "sword fighting" and limited maneuverability (surgical freedom). The newly introduced concept of surgical freedom is very essential in surgical planning and approach selection and can play a role in determining outcome of the procedure, since limited surgical freedom can cause fatigue or limit the extent of lesion resection. In my thesis, we develop a consistent objective methodology to quantify and evaluate surgical freedom, which has been previously evaluated subjectively, and apply this model to the analysis of various endoscopic techniques. This model is crucial for evaluating different endoscopic surgical approaches before they are applied in a clinical setting, for identifying surgical maneuvers that can improve surgical freedom, and for developing endoscopic training simulators that accurately model the surgical freedom of various approaches. Quantifying the extent of endoscopic surgical freedom will also provide developers with valuable data that will help them design improved endoscopes and endoscopic instrumentation. / Dissertation/Thesis / Ph.D. Neuroscience 2014

<div>Experiments were performed on a sharp-nosed 7° half-angle cone at a 6° angle of attack in the Boeing/AFOSR Mach-6 Quiet Tunnel (BAM6QT) to study the stationary crossflow instability and its receptivity to small surface roughness. Heat transfer measurements were obtained using temperature sensitive paint (TSP) and Schmidt Boelter (SB) heat transfer gauges. Great care was taken to obtain repeatable, quantitative measurements from TSP.</div><div></div><div>Consecutive runs were performed at a 0° angle of attack, and the heat transfer measured by the SB was found to drop as the initial model temperature increased, while other initial conditions such as stagnation pressure were held constant. This agreed with calculations done using a similarity solution. It was found that repeatable measurements at a 6° angle of attack could be made if the initial model temperature was controlled and the patch location that was used to calibrate the TSP was picked in a reasonable and consistent manner.</div><div></div><div>The Rod Insertion Method (RIM) roughness, which was used to excite the stationary crossflow instability, was found to be responsible for the appearance of the streaks that were analyzed. The signal-to-noise ratio in the TSP was too low to properly measure the streaks directly downstream of the roughness insert. The heat transfer along the streak experienced linear growth, peaked, and then slightly decayed. It is possible this peak was saturation. The general trend was that the growth of the streaks moved farther upstream as the roughness element height increased, which agreed with past computations and low speed experiments. The growth of the streak also moved farther upstream as the freestream Reynolds number increased. The amplitude of the streaks was calculated by non-dimensionalizing the heat transfer using the laminar theoretical mean-flow solution for a 7° half-angle cone at a 6° angle of attack. The relationship between the amplitude and the non-dimensional roughness height was approximately linear in the growth region of the streaks.</div>

A Laser Inertial Fusion Energy (LIFE) target’s flight through a low Reynolds number and high Mach number regime was analyzed with computational fluid dynamics software. This regime consisted of xenon gas at 1,050 K and approximately 6,670 Pa. Simulations with similar flow conditions were performed over a sphere and compared with experimental data and published correlations for validation purposes. Transient considerations of the developing flow around the target were explored. Simulations of the target at different velocities were used to determine correlations for the drag coefficient and Nusselt number as functions of the Reynolds number. Simulations with different target angles of attack were used to determine the aerodynamic coefficients of drag, lift, Magnus moment, and overturning moment as well as target stability. The drag force, lift force, and overturning moment changed minimally with spin. Above an angle of attack of 15°, the overturning moment would be destabilizing. At angles of attack less than 15°, the overturning moment would tend to decrease the target’s angle of attack, indicating the lack of a need for spin for stability at these small angles. This stabilizing moment would cause the target to move in a mildly damped oscillation about the axis parallel to the free-stream velocity vector through the target’s center of gravity.

The objective of the present research is to investigate unsteady aerodynamic models with state equation representations that are valid up to the high angle of attack regime with the purpose of evaluating them as computationally affordable models that can be used in conjunction with the equations of motion to simulate wing rock. The unsteady aerodynamic models with state equation representations investigated are functional approaches to modeling aerodynamic phenomena, not directly derived from the physical principles of the problem. They are thought to have advantages with respect to the physical modeling methods mainly because of the lower computational cost involved in the calculations. The unsteady aerodynamic multi-axis models with state equation representations investigated in this report assume the decomposition of the airplane into lifting surfaces or panels that have their particular aerodynamic force coefficients modeled as dynamic state-space models. These coefficients are summed up to find the total aircraft force coefficients. The products of the panel force coefficients and their moment arms with reference to a given axis are summed up to find the global aircraft moment coefficients. Two proposed variations of the state space representation of the basic unsteady aerodynamic model are identified using experimental aerodynamic data available in the open literature for slender delta wings, and tested in order to investigate their ability to represent the wing rock phenomenon. The identifications for the second proposed formulation are found to match the experimental data well. The simulations revealed that even though it was constructed with scarce data, the model presented the expected qualitative behavior and that the concept is able to simulate wing rock. / Ph. D.

No description available.

Harmonic oscillatory tests for a fighter aircraft using the Dynamic Plunge--Pitch--Roll model mount at Virginia Tech Stability Wind Tunnel are described. Corresponding data reduction methods are developed on the basis of multirate digital signal processing. Since the model is sting mounted, the frequencies associated with sting vibration are included in balance readings thus a linear filter must be used to extract out the aerodynamic responses. To achieve this, a Finite Impulse Response (FIR) is designed using the Remez exchange algorithm. Based on the reduced data, a state–space model is developed to describe the unsteady aerodynamic characteristics of the aircraft during roll oscillations. For this model, we chose to separate the aircraft into panels and model the local forces and moments. Included in this technique is the introduction of a new state variable, a separation state variable which characterizes the separation for each panel. This new variable is governed by a first order differential equation. Taylor series expansions in terms of the input variables were performed to obtain the aerodynamic coefficients of the model. These derivatives, a form of the stability derivative approach, are not constant but rather quadratic functions of the new state variable. Finally, the concept of the model was expanded to allow for the addition of longitudinal motions. Thus, pitching moments will be identified at the same time as rolling moments. The results show that the goal of modeling coupled longitudinal and lateral–directional characteristics at the same time using the same inputs is feasible. / Master of Science

In this study, the cooling of a simulated blade under increasing pitch between holes was examined. The change in non-dimensional surface temperature, phi, was measured experimentally to quantify this performance loss. This critical quantification of the sensitivity of cooling to pitch between holes has not been studied previously. A range of blowing ratios and angles of attack were tested. Data are presented in terms of the laterally averaged phi, and in terms of the minimum phi, arguably more important from a design perspective. Increasing the pitch 13% produced no measureable change using either parameter. An increase of 26% in pitch produced only a 4% loss in lateral averages, while some hot points dropped by 10%. These small changes are due to compensating effects of increased internal and through-hole convective cooling. A limit to these effects was shown when increasing pitch 53%. While performance loss in the average was still relatively small at 15%, the minimum phi decreased by 27%. Heat flux gauges were used to gather data on the internal and external surface. The internal impingement used in this study represents a more accurate representation of internal cooling for an actual engine part than has been previously studied, providing a starting point for exploring the differences between engine configurations and those generally investigated in the literature. External heat flux measurements were used to measure the ratio of heat flux with and without film cooling. These results call into question the use of the net heat flux reduction parameter, which is commonly used to quantify overall film cooling performance. / text

Baigiamajame magistro darbe atliekamas sklandytuvo atakos ir slydimo kampo matavimo metodų tyrimas. Pirmoje darbo dalyje apžvelgiami atakos kampo matuokliai: virvutė, pritvirtinta ant stiklinio gaubto, atakos – slydimo kampo matuoklis su vėjarodžiu ir Pitoto vamzdelio tipo daviklis. Davikliai palyginami, aprašomi jų privalumai ir trūkumai lyginant su kitais davikliais. Antroje dalyje aprašomi alfa ir beta kampų matavimo metodai: matavimas vamzdelio tipo davikliu ir metodas, kai nenaudojami specialūs atakos, slydimo kampo davikliai. Toliau apžvelgiami veiksniai, turintys įtakos matavimo tikslumui. Pateikiamos kelių vamzdelio tipo daviklių kalibravimo kreivės. Paskutinėje dalyje programa Matlab kuriamas matematinis-dinaminis sklandytuvo modelis. Modeliu, pagal nustatytas sąlygas, skaičiuojami atakos ir slydimo kampai, analizuojami grafikai. / The thesis made the glider’s attack and slip angle measurement methods for the investigation. The first part gives an overview of measuring devices of angle of attack and slip angle: The side string, attached to the side of the canopy, vane mounted AOA sensor, Pitot-tube type sensor. The sensors are compared, describes their advantages and disadvantages compared with other sensors. The second part describes the alpha and beta angle measurement methods: measurement with the tube-type sensor, and the method without using the specific attack, slip angle sensors. The following gives an overview of factors affecting the measurement accuracy. Several tube-type sensor calibration curves are presented. The last part of thesis presents development of mathematical – dynamic model of the glider using Matlab software. The model calculates the angle of attack and slip using set conditions of flight.

Approved for public release; distribution is unlimited. / Ever-increasing demands for accuracy and range in modern warfare have expedited the optimization of projectile design. The crux of projectile design lies in the understanding of its aerodynamic properties early in the design phase. This research first investigated the aerodynamic properties of a standard M549, 155mm projectile. The transonic speed region was the focus of the research as significant aerodynamic variation occurs within this particular region. Aerodynamic data from wind tunnel and range testing was benchmarked against modern aerodynamic prediction programs like ANSYS CFX and Aero-Prediction 09 (AP09). Next, a comparison was made between two types of angle of attack generation methods in ANSYS CFX. The research then focused on controlled tilting of the projectile’s nose to investigate the resulting aerodynamic effects. ANSYS CFX was found to provide better agreement with the experimental data than AP09.