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Guarendi, Andrew NNumerical Investigations of Magnetohydrodynamic Hypersonic Flows
Master of Science, University of Akron, 2013, Mechanical Engineering
Numerical simulations of magnetohydrodynamic (MHD) hypersonic flow are presented for both laminar and turbulent flow over a cylinder and flow entering a scramjet inlet. ANSYS CFX is used to carry out calculations for steady flow at hypersonic speeds (Mach number > 5). The low magnetic Reynolds number (<<1) calculated based on the velocity and length scales in this problem justifies the quasistatic approximation, which assumes negligible effect of velocity on magnetic fields. Therefore the governing equations employed in the simulations are the compressible Navier-Stokes and the energy equations with MHD-related source terms such as Lorentz force and Joule dissipation. Turbulence effects are accounted for when applicable and multiple turbulence models are compared. The results demonstrate the ability of the magnetic field to affect the flowfield, and variables such as location and magnitude of the applied magnetic field are examined. An examination of future work is provided through the implementation of a semi-discrete central scheme in-house code toward the solution of the Orszag-Tang vortex system.

Committee:

Abhilash Chandy, Dr. (Advisor); Scott Sawyer, Dr. (Committee Member); Alex Povitsky, Dr. (Committee Member)

Subjects:

Aerospace Engineering; Engineering; Fluid Dynamics; Mechanical Engineering

Keywords:

Hypersonic; Hypersonic Flow; Flow over a cylinder; Magnetohydrodynamic; MHD; Lorentz; Hypersonic MHD; Numerical Methods; CFD; Computational fluid dynamics; fluid dynamics; Aerospace;

Sagerman, Denton GregoryHypersonic Experimental Aero-thermal Capability Study Through Multilevel Fidelity Computational Fluid Dynamics
Master of Science (M.S.), University of Dayton, 2017, Aerospace Engineering
As true with all hypersonic flight, the ability to quickly and accurately predict the aero-thermodynamic response of an aircraft in the early design phase is important to not only lower cost, but also to lower the computational and experimental time required to test various parameters. The Mach 6 High Reynolds Number Facility at Wright-Patterson Air Force Base in Dayton, Ohio has been non-operational for the past twenty years, but a recent resurgence in the need for accurate hypersonic test facilities has led to the reactivation of the wind tunnel. With its restoration, new capabilities to assess hypersonic aero-thermodynamic effects on bodies in Mach 6 flow have emerged. Therefore, the objective of this research is to determine if obtaining aero-thermal data from the Mach 6 tunnel using temperature sensitive paint (TSP) is a viable option. Surface pressure and temperature readings, from pressure taps and thermocouples installed on the models, as well as TSP wall temperature distributions will be used for comparison with results from computational fluid dynamics (CFD) analysis codes of differing fidelity levels. The comparisons can then be utilized to gain confidence in the ability of the tunnel to capture the aero-thermal response of complex geometries. Three computational codes were used for numerical comparisons: the Configuration Based Aerodynamics (CBAero) tool set, an inviscid panel code with viscous approximation capabilities, Cart3d coupled with Unstructured Langley Approximate Three-Dimensional Convective Heating (UNLATCH) code to approximate viscous effects from the Euler solution, and finally Fun3d, a fully viscous RANS solver. The three tunnel model geometries that will be used for this research are the Reference Flight System model G (RFSG), a Generic Hypersonic Vehicle (GHV), and the Hypersonic International Flight Research Experimentation Program-Flight 1 (HIFiRE-1) payload geometry.

Committee:

Markus Rumpfkeil, Dr. (Committee Chair); Aaron Altman, Dr. (Committee Member); Jose Camberos, Dr. (Committee Member)

Subjects:

Aerospace Engineering

Keywords:

hypersonics; hypersonic experimental testing; aero-thermal; temperature sensitive paint; CFD; hypersonic CFD; tsp

Perkins, Hugh DouglasDevelopment and Demonstration of a Computational Tool for the Analysis of Particle Vitiation Effects in Hypersonic Propulsion Test Facilities
Doctor of Philosophy, Case Western Reserve University, 2009, EMC - Mechanical Engineering

In order to improve the understanding of particle vitiation effects in hypersonic propulsion test facilities, a quasi-one dimensional numerical tool was developed to efficiently model reacting particle-gas flows over a wide range of conditions. Features of this code include gas-phase finite-rate kinetics, a global porous-particle combustion model, mass, momentum and energy interactions between phases, and subsonic and supersonic particle drag and heat transfer models. The basic capabilities of this tool were validated against available data or other validated codes.

To demonstrate the capabilities of the code, and to provide initial insight into the effects of various particle laden flows on ignition, a series of computations were performed for a model hypersonic propulsion test facility and scramjet. Parameters studied were simulated flight Mach number (Mach 5, 6, and 7), particle size (10, 100, and 1000 micron diameters), particle mass fraction (single particle and 1%) and particle material (alumina and graphite). For the alumina particles, it was found that the presence of particles up to 1% mass fraction had very little effect on the gas phase, even though only the 10 micron particles closely followed the gas flow velocity and temperature. With the graphite particles, the 10 micron particles were either quickly quenched, or were quickly consumed, depending on the gas temperature. As the particle size was increased to 100 microns, the particles did not quench, but were still typically consumed within the model test facility. For the 1000 micron particles, combustion was diffusion limited, so particle and gas temperature had little effect on the combustion rate. When the particle mass fraction was increased to 1%, the main change was the addition of significant heat release. In those cases where low graphite reaction rates were observed for single particles, the increase to 1% mass fraction had very little impact.

Hydrogen/vitiated air ignition delay calculations for the 1% mass fraction of graphite particles cases showed significant decreases in ignition delay in cases where higher graphite combustion rates were observed. Further calculations showed that this was due primarily to increased combustor inlet temperature, not the gaseous or solid vitiate species present in the flow.

Committee:

Chih-Jen Sung, PhD (Advisor); James Tien, PhD (Committee Member); Yasuhiro Kamotani, PhD (Committee Member); Steven Izen, PhD (Committee Member)

Subjects:

Engineering; Mechanical Engineering

Keywords:

Gas-Particulate Flows; Hypersonic Wind Tunnel; Graphite Combustion; Flow Vitiation; Scramjet Testing; Hypersonic Propulsion; Particle Combustion

Morris, Seth HendersonQuasi-Transient Calculation of Surface Temperatures on a Reusable Booster System with High Angles of Attack
Master of Science (M.S.), University of Dayton, 2011, Aerospace Engineering
The calculation of a recovery temperature based heat transfer coefficient proves to be sufficiently independent of wall temperature to use in a three dimensional, transient temperature model of a thermal protection system of a reusable booster concept. After a derivation of recovery temperature from the 1st law of thermodynamics, the weak dependence of the recovery temperature based heat transfer coefficient is investigated by 72 Computational Fluid Dynamics (CFD) models at angles of attack ranging from 0° to 90° over a range of Mach numbers, from Mach 2 to 5, and a variety of thermal boundary conditions at the wall, from isothermal to a conductive wall. Then, the heat transfer coefficient is calculated at many steady state CFD solutions for a reusable booster system concept on a given trajectory and applied to a transient Finite Element Analysis (FEA) model of a thermal protection system. Results are presented graphically.

Committee:

Timothy J. Fry, PhD (Committee Chair); José A. Camberos, PhD, PE (Committee Member); John Doty, PhD (Committee Member)

Subjects:

Aerospace Engineering; Atmosphere; Computer Science; Fluid Dynamics; High Temperature Physics; Mechanical Engineering

Keywords:

quasi; transient; heat; transfer; coefficient; hypersonic; CFD; computational; fluid; dynamics; thermodynamic; FEA; steady; state; recovery; temperature; reusable; booster; system; high; speed; computing; super; computer; kinetic; energy

Andrienko, DaniilNon-equilibrium Models for High Temperature Gas Flows
Doctor of Philosophy (PhD), Wright State University, 2014, Engineering PhD
A multi-physical approach is developed for non-equilibrium high enthalpy flows to cover a wide range of interdisciplinary studies. Compressible Navier-Stokes equations are coupled with the chemical model and equation of energy conservation for internal degrees of freedom. The model is capable of accounting for the non-equilibrium dissociation and thermal radiation heat transfer in gas flows. The methodology covers a range of problems from the combustion in subsonic counter-flowing jets to the hypersonic re-entry flows. The present approach attempts to validate the finite rate chemistry mechanisms by modeling the diffusive non-premixed flame of hydrogen at a very low Mach number. The comparison with the artificial compressibility method provides some insights into the compressible Navier-Stokes equations when applied to the flow with nearly constant pressure. High Mach number flow is assessed by a TVD flux splitting scheme applied to model the re-entry of RAM-C II probe. The chemical composition of the air plasma is validated against available experimental and theoretical data. The uncertainty in electron concentration, measured by microwave reflectometers and a Langmuir probe, is examined in the light of chemical reactions, multi-temperature models, boundary conditions and a vibrational-dissociation coupling model. The heating rates are reported on a wide range of trajectory points. A novel approach is used to improve the accuracy of the radiation transfer model based on the ray tracing method by introducing Gauss-Lobatto quadrature and space partition algorithms to the nearest neighbor search. The accuracy of the method is analyzed for hypersonic flows and for laser-supported combustion waves. The approach revealed an increased efficiency of two orders of magnitude when solving the radiation transfer energy along the line of sight. Finally, the concept of the view factor is applied to assess the radiation transfer in high temperature absorbing gas flows. This concept was verified against the ray tracing method and demonstrated several important advantages. The semi-analytical solution for the radiation flux density in axisymmetric geometry is obtained in terms of elliptic integrals. The cost of the radiation transfer model based on the view factor approach in weakly absorbing media is comparable with the cost of the tangent slab approximation and achieves the asymptotic accuracy of the ray tracing method.

Committee:

George Huang, Ph.D. (Advisor); Joseph Shang, Ph.D. (Advisor); James Menart, Ph.D. (Committee Member); William Bailey, Ph.D. (Committee Member); Paul King, Ph.D. (Committee Member); Viswanath Katta, Ph.D. (Committee Member)

Subjects:

Aerospace Engineering; Fluid Dynamics; Mechanical Engineering

Keywords:

Radiation transfer, Finite-rate chemistry, air chemistry, reentry, hypersonic flow, numerical models, gasdynamics, ray tracing method, zonal method, hydrogen combustion

LeVett, Marshall AllanParallel Time-Marching for Fluid-Thermal-Structural Interactions
Master of Science, The Ohio State University, 2016, Aero/Astro Engineering
This thesis evaluates parallel-in-time methods for application upon fluid-thermal-structural interactions, particularly those that could be seen on a reusable ultra-high-speed flight vehicle. The United States Air Force is interested in this class of vehicles due to their potential for long-range, high-speed strikes. However, there are design hurdles for a vehicle in this class because of extreme aerodynamic heating and loading. The coupled, nonlinear response of the structure, fluid, and temperature leads to the need for computationally intensive time-accurate response. Parallel-in-time methods offer potential to speedup the time-to-solution of such complicated time-dependent problems by leveraging ubiquitous parallel computing resources in an additional branch of parallelism beyond parallel-in-space methods. The stability and speedup of multiple parallel-in-time methods are first discussed. The numerical stability of a common parallel-in-time method is evaluated on a linear test function, showing instability for problems with imaginary dominated eigenvalues which appear frequently in structural dynamics. Instability is also linked to invariant breaking behavior of parallel-in-time methods. A particular parallel-in-time method -- developed previously for nonlinear structural systems -- is adapted for application to fluid-thermal-structural analysis. The resulting method is applied to a panel reprentative of what could be seen on a future high-speed-flight vehicle. The method is first examined for predicting the aeroelastic response of a panel in supersonic flow, and the result converges to a serial solution in both aeroelastically stable and unstable regimes. Theoretical speedups are found to be between 1.6 to 2.2 over a serial solution. Next, the scheme is assessed for computing the transient fluid-thermal-structural response of a panel in supersonic flow. The result converges to the serial solution until an aerothermoelastically induced snapthrough, after which the scheme is unstable. Theoretical speedup in the stable region of this situation is found to be 5.5 over a serial solution.

Committee:

Jack McNamara (Advisor); Soghrati Soheil (Committee Member)

Subjects:

Aerospace Engineering

Keywords:

parareal; parallel; in; time; fluid; thermal; structural; interactions; multiphysics; hypersonic; reusable; modeling; simulation

Atkinson, Michael D.Control of Hypersonic High Angle-Of-Attack Re-Entry Flow Using a Semi-Empirical Plasma Actuator Model
Doctor of Philosophy (Ph.D.), University of Dayton, 2012, Aerospace Engineering

The aim of this dissertation was to explore the possibility of using flow control to stabilize re-entry flight at very high angle-of-attack. This was carried out in three steps: 1) study the structure of representative high angle-of-attack re-entry flows; 2) develop a semi-empirical plasma actuator model that can be applied to control high angle-of-attack re-entry flows; 3) application of the plasma actuator model to study the control of representative re-entry flows. The calculations include viscous and thermochemical non-equilibrium effects, and a high-fidelity physical model to resolve complex flow structure.

The contribution of this dissertation was to provide a detailed description of hypersonic viscous flow around blunt-nosed elliptical cone at very high angle-of-attack. High-fidelity, thermochemical non-equilibrium numerical solutions of high angle-of-attack re-entry flows were not published prior to this research, and thus this research can provide a foundation to calculate, analyze, and describe very high angle-of-attack hypersonic re-entry flows.

Paramount to this dissertation was the development of a new phenomenological MHD plasma actuator model. A semi-empirical actuator model was developed by adding source terms to the momentum equation, vibrational energy equation, and total energy equation, employing an exponential decay function based on the formulations of Kalra et al. and Poggie. This new plasma actuator model was extended from Poggie's model to include thermochemical non-equilibrium effects and expanded from Kalra's et al. two-dimensional model to include three-dimensional effects. Development, validation, and calibration of the plasma actuator model was based on a qualitative comparison to the experiment of Kalra et al. on manipulating turbulent shock-wave/bounday layer interaction using plasma actuators. The effect of the plasma actuators on turbulent shock-wave/boundary-layer interaction was simulated numerically and a detailed description of the complex flow structure with and without actuation was provided.

Finally, application of the plasma actuators to control the complex flow structure of high angle-of-attack re-entry flight vehicles was investigated. To the best of the author's knowledge, no prior research on high angle-of-attack re-entry vehicle control using plasma actuators has been published. Lastly, this dissertation serves as a foundation to compute, analyze, and control complex flow generated around re-entry vehicles at high angle-of-attack.

Committee:

José A. Camberos, PhD (Committee Chair); Jonathan Poggie, PhD (Committee Co-Chair); Aaron M. Altman, PhD (Committee Member); Youssef N. Raffoul, PhD (Committee Member)

Subjects:

Aerospace Engineering

Keywords:

Hypersonic, Reentry; CFD; Plasma actuators; Flow Control; High angle of attack

Gunbatar, YakupNonlinear Adaptive Control and Guidance for Unstart Recovery for a Generic Hypersonic Vehicle
Doctor of Philosophy, The Ohio State University, 2014, Electrical and Computer Engineering
This work presents the development of an integrated flight controller for a generic model of a hypersonic air-breathing vehicle. The flight control architecture comprises a guidance and trajectory planning module and a nonlinear inner-loop adaptive controller. The emphasis of the controller design is on achieving stable tracking of suitable reference trajectories in the presence of a specific engine fault (inlet unstart), in which sudden and drastic changes in the vehicle aerodynamics and engine performance occur. First, the equations of motion of the vehicle for a rigid body model, taking the rotation of the Earth into account, is provided. Aerodynamic forces and moments and engine data are provided in lookup-table format. This comprehensive model is used for simulations and verification of the control strategies. Then, a simplified control-oriented model is developed for the purpose of control design and stability analysis. The design of the guidance and nonlinear adaptive control algorithms is first carried out on a longitudinal version of the vehicle dynamics. The design is verified in a simulation study aiming at testing the robustness of the inner-loop controller under significant model uncertainty and engine failures. At the same time, the guidance system provides reference trajectories to maximize the vehicle's endurance, which is cast as an optimal control problem. The design is then extended to tackle the significantly more challenging case of the 6-degree-of-freedom (6-DOF) vehicle dynamics. For the full 6-DOF case, the adaptive nonlinear flight controller is tested on more challenging maneuvers, where values of the flight path and bank angles exceed the nominal range defined for the vehicle. Simulation studies show stable operation of the closed-loop system in nominal operating conditions, unstart conditions, and during transition from sustained scramjet propulsion to engine failure mode.

Committee:

Andrea Serrani, Prof. (Advisor); Umit Ozguner, Prof. (Committee Member); Zhang Wei, Prof. (Committee Member)

Subjects:

Aerospace Engineering; Computer Engineering; Electrical Engineering; Engineering

Keywords:

Hypersonic; unstart; adaptive control; nonlinear control; adaptive backstepping; trajectory optimization; optimal; tuning functions; longitudinal; 6-DOF; 3-DOF; Earth rotation; Control design model; coordinated turn; endurance; Equations of motion

Sigthorsson, David O.Control-Oriented Modeling and Output Feedback Control of Hypersonic Air-Breathing Vehicles
Doctor of Philosophy, The Ohio State University, 2008, Electrical and Computer Engineering
Hypersonic air-breathing vehicles are a promising and cost-efficient technology for launching low-earth-orbit satellites and providing rapid global-response capabilities. Modeling and control of such vehicles has been an active subject of research in recent years. A first-principle, physics-based model (FPM) of the vehicle's longitudinal dynamics has been developed at the Air Force Research Laboratory, and made available to the academic community for control systems design. This model, while suitable for simulation, is intractable for model-based control, thus requiring further control-oriented modeling. A typical control objective is to track a velocity and altitude reference while maintaining physical feasibility of the control input and the state. Two control strategies are presented in this work. The first is a linear time invariant (LTI) design based on a novel formulation of a robust servo-mechanism using singular perturbation arguments. This approach does not rely on state reconstruction but does require an analysis of a family of linearized models from the FPM. The second design relies on reduced-complexity modeling of the FPM. Intractable expressions of the forces and moment in the FPM are replaced with a curve-fit model (CFM). The CFM is expressed as a linear parameter varying (LPV) system, where the scheduling variables depend on the system output. A novel LPV regulator design methodology is developed, which explicitly addresses the case of over-actuated models (i.e., models with more inputs than performance outputs). This is a non-trivial extension of the analysis and design of output regulators for LTI systems. The LPV regulator separates the control problem into a steady-state controller and a stabilizing controller. The steady-state controller produces a non-unique approximate steady-state using receding horizon constrained optimization, while the stabilizer renders the steady-state attractive. The steady-state controller represents an approach to addressing over-actuated LPV systems, alternative to static or dynamic control allocation, or standard optimal control. The stabilizer design utilizes the LPV separation principle to decompose the problem into state feedback and LPV reduced order observer design. Both approaches are applied to the FPM in simulation and their merits and drawbacks discussed.

Committee:

Andrea Serrani, PhD (Advisor); Stephen Yurkovich, PhD (Committee Member); Kevin Passino, PhD (Committee Member)

Subjects:

Electrical Engineering; Engineering

Keywords:

Electrical Engineering; Control Design; LPV Modeling; LPV Control; MIMO systems; Hypersonic Air-Breathing Vehicles; Non-Linear Modeling; Control Oriented Modeling; System ID; Over-Actuated Systems; LPV Regulator; Robust Control; Output-Feedback Control

Parker, Grant HoustonPyrolytic Decomposition of Synthetic Paraffinic Kerosene Fuel Compared to JP-7 and JP-8 Aviation Fuels
Master of Science (M.S.), University of Dayton, 2013, Chemical Engineering
Every generation of advanced military aircraft fly higher and faster than previous generations. With these leaps in performance, aircrafts develop enormous heat loads which can exceed aircraft material limitations. To relieve these heat loads, aircraft can utilize the endothermic heat sink capacity of jet fuel realized through pyrolytic decomposition. Improved understanding of the effect of fuel chemical composition on supercritical pyrolytic reactivity under conditions relevant to advanced aircraft operation can assist with the successful development of viable cooling methodologies. The goal of the current study was to compare the pyrolytic reactivity, primary decomposition products, and global reaction rates of fuels with varying chemical composition. A flowing reactor system was used to explore the pyrolytic chemistry of a Synthetic Paraffinic Kerosene (SPK) and specification jet fuels JP-8 and JP-7. The SPK was comprised solely of iso- and n-paraffins, with negligible cycloparaffin and aromatic content, while the specification fuels had chemical compositions consistent with typical petroleum-derived fuels. The pyrolytic studies were performed using stainless steel tube reactors which were 37.5 cm long and 0.5 mm inside diameter, with inlet flow rates of 0.2 to 0.6 mL/min at a pressure of 3.54 MPa. External reactor wall temperatures ranged from 500C to 650C. The liquid to gas conversion by mass was used as a metric for evaluating the pyrolytic reactivity due to the complex multicomponent composition of the test fuels. SPK averaged 45% higher conversion than JP-7 and 75% higher conversion than JP-8 at each respective temperature. All fuels followed similar reactivity trends with respect to controlling reaction chemistry, such as such as decomposition of long chain nparaffins, olefin formation, cycloparaffins formation, aromatic formation, and gas (e.g.,low molecular weight compound) production. Characterization of the relative reactivity of the fuels was performed by assuming the fuels decomposed via a first order, irreversible reaction pathway with respect to the gravimetric liquid to gas conversion. The calculated reaction rates and temperature data were used to develop Arrhenius plots which yielded the following kinetic perimeters: SPK--pre-exponential (A) factor of 2.3 x 10^12 s-1 and activation energy (Ea) of 223 kcal/mol, JP-7--A of 2.1 x 10^12 s-1 and Ea of 226 kcal/mol, and JP-8--A of 4.6 x 10^12 s-1 and Ea of 235 kcal/mol. These parameters can be used to estimate the initial reactivity and decomposition of these fuels under endothermic conditions. SPK fuels are more pyrolytically reactive compared to JP-7 and JP-8 using the liquid to gas conversion metric due to the variation in the neat chemical compositions. The mildly branched paraffins of the SPK with negligible cycloparaffins and aromatics, which can act as hydrogen donors reducing propagation rate, limited the reaction pathways resulting in a high liquid to gas conversion. JP-7 and JP-8 had a lower liquid to gas conversion due to the significantly higher initial concentrations of cycloparaffins and aromatics, thereby enabling these fuels to participate in a greater number of hydrogen donor reactions which lowers the extent of propagation reactions. Implications of these results can vary depending on the heat sink design and endothermic fuel cooling strategy. The propensity of SPK to react at lower temperature can enable SPK fuels to more readily reach the endothermic heating value. Unfortunately, a fuel with a higher pyrolytic reaction rate can also produce carbon deposition more readily. Further development into the design of a hypersonic heat exchanger system and the determination of the acceptable amount of liquid to gas conversion will dictate the optimal endothermic fuel.

Committee:

Matthew DeWitt (Advisor); Kevin Myers (Committee Member); Steven Zabarnick (Committee Member); Richard Striebich (Committee Member)

Subjects:

Aerospace Engineering; Alternative Energy; Chemical Engineering

Keywords:

pyrolysis, endothermic jet fuels, synthetic paraffinic kerosene, hypersonic, endothermic heat sink, pyrolytic decomposition, JP-7, fischer tropsch, FT

Fiorentini, LisaNonlinear Adaptive Controller Design For Air-breathing Hypersonic Vehicles
Doctor of Philosophy, The Ohio State University, 2010, Electrical and Computer Engineering
This dissertation presents the design of two nonlinear robust controllers for an air-breathing hypersonic vehicle model capable of providing stable tracking of velocity and altitude (or flight-path angle) reference trajectories. To overcome the analytical intractability of a dynamical model derived from first principles, a simplified control-oriented model is used for control design. The control-oriented model retains the most important features of the model from which it was derived, including the non-minimum phase characteristic of the flight-path angle dynamics and strong couplings between the engine and flight dynamics. The first control design considers as control inputs the fuel equivalence ratio and the elevator and canard deflections. A combination of nonlinear sequential loop-closure and adaptive dynamic inversion has been adopted for the design of a dynamic state-feedback controller. An important contribution given by this work is the complete characterization of the internal dynamics of the model has been derived for Lyapunov-based stability analysis of the closed-loop system, which includes the structural dynamics. The results obtained address the issue of stability robustness with respect to both parametric model uncertainty, which naturally arises in adopting reduced-complexity models for control design, and dynamic perturbations due to the flexible dynamics. In the second control design a first step has been taken in extending those results in the case in which only two control inputs are available, namely the fuel equivalence ratio and the elevator deflection. The extension of these results to this new framework is not trivial since several issues arise. First of all, the vehicle dynamics are characterized by exponentially unstable zero-dynamics when longitudinal velocity and flight-path angle are selected as regulated output. This non-minimum phase behavior arises as a consequence of elevator-to-lift coupling. In the previous design the canard was strategically used to adaptively decouple lift from elevator command, thus rendering the system minimum phase. Moreover, the canard input was also employed to enforce the equilibrium at the desired trim condition and to provide a supplementary stabilizing action. As a result, when this control input is not assumed to be available, the fact that the system needs to be augmented with an integrator (to reconstruct the desired equilibrium) and the non-minimum phase behavior have a strong impact on the control design. In these preliminary results the flexible effects are not taken into account in the stability analysis but are considered as a perturbation and included in the simulation model. The approach considered utilizes a combination of adaptive and robust design methods based on both classical and recently developed nonlinear design tools. As a result, the issue of robustness with respect to parameter uncertainties is addressed also in this control design. Simulation results on the full nonlinear model show the effectiveness of both controllers.

Committee:

Andrea Serrani (Advisor); Stephen Yurkovich (Committee Member); Kevin Passino (Committee Member); Scott Gaudi (Committee Member)

Subjects:

Electrical Engineering

Keywords:

nonlinear control; adaptive control; aircraft control; hypersonic vehicles

Yentsch, Robert JThree-Dimensional Shock-Boundary Layer Interactions in Simulations of HIFiRE-1 and HIFiRE-2
Doctor of Philosophy, The Ohio State University, 2013, Mechanical Engineering
A series of high-fidelity, three-dimensional simulations has been performed to investigate hypersonic phenomena encountered in the HIFiRE Flight 1 and Flight 2 experiments. The investigation of HIFiRE-1highlights the performance of turbulence modeling in realistic hypersonic flight vehicles subject to laminar-to-turbulent boundary layer transition and geometry induced adverse pressure gradient separated shock boundary layer interactions (corner flows) influenced by three-dimensional effects. Comparisons with flight test data indicate that the performance of the turbulence model is dependent on the flow condition and the variable under examination. The surface pressure trends are reproduced in all cases, and predictions for the axial separation location is generally within 20% of the separated region length. For the lower Mach number cases, the surface pressure is predicted better at the lower Reynolds number case. Heat transfer predictions on the cone are good, although the use of empirically specified laminar-to-turbulent transition is necessary. The best comparison in heat transfer rates at the flare are observed at the highest Mach number. Overall, the results suggest that the K-ω turbulence model used in this study can be used for flight test prediction, though such uses must be done with care. The primary focus of this dissertation is on the HIFiRE-2 scramjet, specifically, the transient process of dual-to-scramjet mode-transition. For this geometry, the role of the primary fuel injectors in scramjet-mode operation is very important. The barrel shocks from the jet-in-crossflow interaction serves as a flameholder, allowing upstream propagation of pressure rise from the combustion in the cavity into the isolator. It is also shown that the presence of inlet shocks in theisolator can profoundly change the flow, a fact which must be considered in ground testing. A mode-transition event is present which demarcates dual-mode operation from scramjet-mode operation. This event occurs when the barrel shock from a primary injector collapses and the flow reattaches to the wall, profoundly changing the kinematics of the flow. Hysteresis is observed between the steady-state simulations and the mode-transition simulations, and the root cause is traced to an interaction between the corner separation and the barrel shock from the outboard injectors. Generation of turbulence primarily occurs at the shear layer between reverse flow in the separated region and the supersonic core flow. The overall flow, particularly in the cavity, is highly threedimensional, with strong, first-order interactions between the primary injectors and the high speed isolator flow. Capturing these interactions is critical to predicting the transient phenomena.

Committee:

Datta Gaitonde (Advisor); Michael Brown (Committee Member); Jen-Ping Chen (Committee Member); Jack McNamara (Committee Member); Mei Zhuang (Committee Member)

Subjects:

Aerospace Engineering

Keywords:

HIFiRE; mode-transition; mode transition; scramjet, hypersonic airbreathing propulsion; shock boundary layer interactions; computational fluid dynamics; CFD; RANS; dual-mode; scramjet-mode;

Vick, Tyler JGeometry Modeling and Adaptive Control of Air-Breathing Hypersonic Vehicles
MS, University of Cincinnati, 2014, Engineering and Applied Science: Aerospace Engineering
Air-breathing hypersonic vehicles have the potential to provide global reach and affordable access to space. Recent technological advancements have made scramjet-powered flight achievable, as evidenced by the successes of the X-43A and X-51A flight test programs over the last decade. Air-breathing hypersonic vehicles present unique modeling and control challenges in large part due to the fact that scramjet propulsion systems are highly integrated into the airframe, resulting in strongly coupled and often unstable dynamics. Additionally, the extreme flight conditions and inability to test fully integrated vehicle systems larger than X-51 before flight leads to inherent uncertainty in hypersonic flight. This thesis presents a means to design vehicle geometries, simulate vehicle dynamics, and develop and analyze control systems for hypersonic vehicles. First, a software tool for generating three-dimensional watertight vehicle surface meshes from simple design parameters is developed. These surface meshes are compatible with existing vehicle analysis tools, with which databases of aerodynamic and propulsive forces and moments can be constructed. A six-degree-of-freedom nonlinear dynamics simulation model which incorporates this data is presented. Inner-loop longitudinal and lateral control systems are designed and analyzed utilizing the simulation model. The first is an output feedback proportional-integral linear controller designed using linear quadratic regulator techniques. The second is a model reference adaptive controller (MRAC) which augments this baseline linear controller with an adaptive element. The performance and robustness of each controller are analyzed through simulated time responses to angle-of-attack and bank angle commands, while various uncertainties are introduced. The MRAC architecture enables the controller to adapt in a nonlinear fashion to deviations from the desired response, allowing for improved tracking performance, stability, and robustness.

Committee:

Kelly Cohen, Ph.D. (Committee Chair); Michael Bolender, Ph.D. (Committee Member); Elad Kivelevitch, Ph.D. (Committee Member)

Subjects:

Aerospace Materials

Keywords:

Hypersonic Vehicle;Adaptive Control;Geometry Modeling;Modeling and Simulation;Model Reference;Linear Quadratic Regulator

Smith, Joshua GabrielLoosely Coupled Hypersonic Airflow Simulation over a Thermally Deforming Panel with Applications for a POD Reduced Order Model
Master of Science, Miami University, 2017, Mechanical & Manufacturing Engineering
Predicting surface panel buckling caused by viscous heating around hypersonic vehicles is a complex problem as the interaction between the airflow and the structure results in a physically coupled system that is computationally expensive to solve with traditional computational fluid dynamics (CFD). Reduced order modeling (ROM) via proper orthogonal decomposition (POD) offers improved efficiency in simulating this system over traditional CFD, but must be formed from an existing data set. This study used loosely coupled fluid-thermal-structural simulations to examine Mach 4.24 flow over a deforming stainless steel panel. The methodology and domain from a previous study were used and expanded. ANSYS Fluent CFD software was used to solve the flow field for steady state and transient scenarios. Abaqus finite element analysis (FEA) software was used to examine the panel’s response to applied temperature profiles. Results were compared to the previous study and oblique shock theory for validation. It was found that altering the interval of data transfer between systems could be optimized to balance accuracy and efficiency. Applications associated with creating a POD ROM from the simulation data were also considered.

Committee:

Edgar Caraballo, Dr. (Advisor); Amit Shukla, Dr. (Committee Member); Andrew Sommers, Dr. (Committee Member)

Subjects:

Aerospace Engineering; Mechanical Engineering

Keywords:

hypersonic; buckling; loosely coupled simulation; fluid-thermal interaction

Benyo, Theresa L.Analytical and Computational Investigations of a Magnetohydrodynamic (MHD) Energy-Bypass System for Supersonic Turbojet Engines to Enable Hypersonic Flight
PHD, Kent State University, 2013, College of Arts and Sciences / Department of Physics
Historically, the National Aeronautics and Space Administration (NASA) has used rocket-powered vehicles as launch vehicles for access to space. A familiar example is the Space Shuttle launch system. These vehicles carry both fuel and oxidizer onboard. If an external oxidizer (such as the Earth's atmosphere) is utilized, the need to carry an onboard oxidizer is eliminated, and future launch vehicles could carry a larger payload into orbit at a fraction of the total fuel expenditure. For this reason, NASA is currently researching the use of air-breathing engines to power the first stage of two-stage-to-orbit hypersonic launch systems. Removing the need to carry an onboard oxidizer leads also to reductions in total vehicle weight at liftoff. This in turn reduces the total mass of propellant required, and thus decreases the cost of carrying a specific payload into orbit or beyond. However, achieving hypersonic flight with air-breathing jet engines has several technical challenges. These challenges, such as the mode transition from supersonic to hypersonic engine operation, are under study in NASA's Fundamental Aeronautics Program. One propulsion concept that is being explored is a magnetohydrodynamic (MHD) energy- bypass generator coupled with an off-the-shelf turbojet/turbofan. It is anticipated that this engine will be capable of operation from takeoff to Mach 7 in a single flowpath without mode transition. The MHD energy bypass consists of an MHD generator placed directly upstream of the engine, and converts a portion of the enthalpy of the inlet flow through the engine into electrical current. This reduction in flow enthalpy corresponds to a reduced Mach number at the turbojet inlet so that the engine stays within its design constraints. Furthermore, the generated electrical current may then be used to power aircraft systems or an MHD accelerator positioned downstream of the turbojet. The MHD accelerator operates in reverse of the MHD generator, re-accelerating the exhaust flow from the engine by converting electrical current back into flow enthalpy to increase thrust. Though there has been considerable research into the use of MHD generators to produce electricity for industrial power plants, interest in the technology for flight-weight aerospace applications has developed only recently. In this research, electromagnetic fields coupled with weakly ionzed gases to slow hypersonic airflow were investigated within the confines of an MHD energy-bypass system with the goal of showing that it is possible for an air-breathing engine to transition from takeoff to Mach 7 without carrying a rocket propulsion system along with it. The MHD energy-bypass system was modeled for use on a supersonic turbojet engine. The model included all components envisioned for an MHD energy-bypass system; two preionizers, an MHD generator, and an MHD accelerator. A thermodynamic cycle analysis of the hypothesized MHD energy-bypass system on an existing supersonic turbojet engine was completed. In addition, a detailed thermodynamic, plasmadynamic, and electromagnetic analysis was combined to offer a single, comprehensive model to describe more fully the proper plasma flows and magnetic fields required for successful operation of the MHD energy bypass system. The unique contribution of this research involved modeling the current density, temperature, velocity, pressure, electric field, Hall parameter, and electrical power throughout an annular MHD generator and an annular MHD accelerator taking into account an external magnetic field within a moving flow field, collisions of electrons with neutral particles in an ionized flow field, and collisions of ions with neutral particles in an ionized flow field (ion slip). In previous research, the ion slip term has not been considered. Detailed thermodynamic cycle analysis of an annular MHD generator and an annular MHD accelerator revealed that including the ion slip term to the generalized Ohm's Law decreased the needed magnetic fields and conductivity levels as compared to previous research. For the MHD generator, the needed magnetic fields decreased from 5 T to 3 T for all flight speeds studied (Mach 7, 5, and 3). The conductivity levels required for the ionized airflow within the MHD generator at 3 T decreased from 11 mhos/m to 9 mhos/m for a flight speed of Mach 7 and remained the same for Mach 5 and 3. For the MHD accelerator, the needed magnetic fields decreased from 5 T to 3 T for flight speeds of Mach 7 and 5, and decreased from 3 T to 1.5 T for a flight speed of Mach 3. The conductivity levels required for the ionized airflow within the MHD accelerator (at 3 T) decreased from 2.6 mhos/m to 1.1 mhos/m for a flight speed of Mach 7 and remained the same for Mach 5 and 3. The MHD energy-bypass system model showed that it is possible to expand the operating range of a supersonic jet engine from a maximum of Mach 3.5 to a maximum of Mach 7. The inclusion of ion slip within the analysis further showed that it is possible to 'drive&apos; this system with maximum magnetic fields of 3 T and with maximum conductivity levels of 11 mhos/m. These operating parameters better the previous findings of 5 T and 10 mhos/m, and reveal that taking into account collisions between ions and neutral particles within a weakly ionized flow provides a more realistic model with added benefits of lower magnetic fields and conductivity levels especially at the higher Mach numbers.

Committee:

David Allendar, PhD (Committee Co-Chair); Isaiah Blankson, PhD (Committee Co-Chair); John Portman, PhD (Committee Member); Mark Manley, PhD (Committee Member); John West, PhD (Committee Member); Jonathan Maletic, PhD (Committee Member)

Subjects:

Aerospace Engineering; Electromagnetics; Theoretical Physics

Keywords:

magnetohydrodynamics; energy bypass; ion slip; supersonic jet propulsion; hypersonic flight; air-breathing jet engine; weakly ionized gases; plasma flows; thermodynamic cycle analysis

Mullenix, Nathan JoelFully Coupled Model for High-Temperature Ablation and a Reative-Riemann Solver for its Solution
Doctor of Philosophy, University of Akron, 2010, Mechanical Engineering
Ablation is a process of rapid material removal from a solid surface by chemical reactions, sublimation and other erosive processes. Ablation absorbs large quantities of heat, which makes it a desirable process for Thermal Protection Systems (TPS) used on aerospace vehicles that encounter severe thermal environments, and has been in use since the beginning of the Space Age. The ablation process consists of several coupled sub-processes including gas dynamics, heat, and ablative mechanisms at the surface. Experimental techniques can reproduce some but not all of the parameters of the flight environment, and for this reason numerical modeling has been undertaken to provide greater understanding. The past state of the art has involved modeling only some of the sub-processes with simplifications of the others, which tended to incorrectly predict the thickness of TPS. The computational resources necessary to model all of the sub-processes in a coupled manner have become available. A tightly coupled mathematical model for the non-charring ablation problem is developed within this dissertation. The Reynolds Transport Theorem (RTT) is used to derive a set of governing equations that takes into account the movement of the ablating surface and the resulting mass transfer. A set of numerical methods has been developed and/or modified from existing forms. For example, finite volume discretization schemes are modified to allow for arbitrary composition (solid or gas) within a fixed body-fitted computational grid, and a new method called a reactive-Riemann solver is derived to solve the mass transfer across the ablating surface and its movement. Finally, serial and parallel algorithms are developed for the numerical methods. Individual components of the model and numerical methods are validated against standard test cases. The full solver is used for ablation problems across a range of free-stream and surface conditions, and these results are shown to agree with experimental data. The effect of surface defects on the local ablation rate and process are shown for the first time: a localized defect can greatly enhance the local ablation rate and create a region of sublimation dominated ablation even if the rest of the surface is ablating primarily via oxidation.

Committee:

Alex Povitsky, Dr. (Advisor); Minel Braun (Committee Member); Scott Sawyer (Committee Member); S. I. Hariharan (Committee Member); Gerald Young (Committee Member)

Subjects:

Aerospace Materials; Engineering; Mechanical Engineering

Keywords:

ablation; tightly coupled solver; finite volume scheme; CFD; graphite; oxidation; nitridation; sublimation; hypersonic flight; TPS

Henderson, Sean JamesStudy of the Issues of Computational Aerothermodynamics Using a Riemann Solver
Doctor of Philosophy (PhD), Wright State University, 2008, Engineering PhD

This work is part of a project to more accurately model hypersonicflow. A number of issues in hypersonic flow are addressed.

The first issue addressed is that of air properties at increased temperatures. In particular the thermodynamic and transport properties of chemical equilibrium air are found for temperatures up to 30,000 K for a pressure range from 1x10-4 to 100 atm. This work provides properties at slightly higher temperatures for the lower pressure region than can be found in the literature. This work also covers adding equilibrium air chemistry to the computational fluid dynamics computer code known as AVUS.

The second issue addressed is commonly referred to as the carbuncle phenomenon. The carbuncle phenomenon is a numerical instability that affects the capturing of strong shocks when using a Riemann solver with low numerical dissipation. The carbuncle phenomenon manifests itself in the inability to compute uniform flow conditions downstream of a normal or nearly normal shock. Prior work has been done in this area to accurately capture strong shocks; and great progress has been made in reducing the effects of the carbuncle phenomenon. Even with these improvements the heat transfer profiles in the stagnation region still show some distortion from small upstream perturbations convected downstream to the wall. It has been determined that the grid quality in the region of the shock is a major factor in the inability of Riemann solvers to accurately capture the flow in the stagnation region. For this reason this work performs a grid study and makes recommendations as to what types of structured grids should be used to accurately capture strong shocks and predict heat transfer profiles at the body surface. This grid study shows that some types of grids suffer more than others from the carbuncle problem. The reason for this is the numerical dissipation that is introduced from the numerical routine. This work shows that grid aspect ratio and the alignment of the grid to the flow can be used to reduce the effects of the carbuncle phenomenon. This work also shows that another mechanism for the carbuncle phenomenon is the alignment of the grids with the shock. The heat transfer profile cannot be properly captured if the grid is not aligned well with the shock.

The third issue addressed in this work is the domain of applicability of the perfect gas model, the equilibrium air model, the nonequilibrium air model, and the thermo-chemical nonequilibrium air model. A computational study is carried out using AVUS to determine the regions of applicability of these air models for a blunt body at various velocities and altitudes. This type of altitude-velocity plot has already been produced by previous researchers, but the dividing lines between the different gas models were found using residence times. This work looks at temperature and heat transfer profiles for a blunt body in a high speed air flow to determine the dividing lines between the regions of applicability of the different air models. Unlike the previous work, this work provides specific error values for using a given model in a certain flight regime. It is found that the dividing lines between chemical equilibrium and chemical nonequilibrium have two dips in the curve that were not shown by previous researchers. These dips correspond to regions where O2 and N2 strongly dissociate.

Committee:

James Menart, Ph.D. (Advisor); Mitch Wolff, Ph.D. (Committee Member); Joseph Shang, Ph.D. (Committee Member); Raymond Maple, Ph.D. (Committee Member); James Miller, Ph.D. (Committee Member)

Subjects:

Fluid Dynamics; Mechanical Engineering

Keywords:

Carbuncle Phenomenon; Riemann Solvers, Hypersonic Flows

Miller, Brent AdamLoosely Coupled Time Integration of Fluid-Thermal-Structural Interactions in Hypersonic Flows
Doctor of Philosophy, The Ohio State University, 2015, Aero/Astro Engineering
The development of reusable hypersonic cruise vehicles requires analysis capability that can capture the coupled, highly-nonlinear interactions between the fluid flow, structural mechanics, and heat transfer. This analysis must also be performed over significant portions of the flight trajectory due to the long-term thermal evolution of the vehicle. The fluid and structural physics operate at significantly smaller time scales than the thermal evolution, requiring time marching that can capture the small time scales for time records that encapsulate the longer time scale of the thermal response. This leads to extreme computational times and motivates research that seeks to maximize efficiency of the time integrations for the coupled problem. The goal of this dissertation is to develop time integration procedures that significantly improve computational efficiency while also maintaining time accuracy and stability for fluid-thermal-structural analysis. This is achieved using carefully designed loosely coupled schemes for the fluid, thermal, and structural solvers. Here, different time integrators are used for the solvers of each physical field, and boundary conditions are exchanged at most once per time step. Coupling schemes for both time-accurate and quasi-steady flow models are considered. Computational efficiency and time accuracy are improved through the use of both extrapolating predictors and interpolation during the exchange of boundary conditions; the latter of which enables the use of different sized time steps between the solvers, known as subcycling. The developed coupling procedures are compared to several other schemes, including a basic one that does not use the predictors, and a subiteration-based strongly coupled scheme. Response predictions of multiple configurations of a panel in two dimensional high-speed flow are performed. Using second order implicit time integrators for the individual solvers, the predictor and strongly coupled schemes are demonstrated to retain the second order accuracy with and without subcycling, while the others reduced to first order. Simulations of two panel configurations are performed to investigate the performance of the coupling schemes in predicting stable and dynamically unstable responses. From the stable response analysis, the predictor-based schemes are found to be the least computationally expensive compared to the strongly coupled and basic loosely coupled schemes. In the second configuration, the panel is predicted to undergo snap-through and ultimately a dynamically unstable limit cycle response with each scheme. The predictor schemes are shown to provide a 3-10 times reduction in computational cost compared to the other schemes. Finally, a 30 second response of the panel with a flutter instability using time-accurate CFD is performed and compared to the response using quasi-steady surrogate and analytical aerothermodynamic models. The unsteady CFD and surrogate based responses have excellent agreement throughout the response, with differences under 1%. The time to flutter predicted from the analytical models is 15% higher than that predicted by the CFD and surrogate models. However, the flutter response shows good agreement between the three responses, indicating that the quasi-steady flow assumption can accurately capture the coupled dynamics.

Committee:

Jack McNamara (Advisor); Thomas Eason, III (Committee Member); Datta Gaitonde (Committee Member); Sandip Mazumder (Committee Member); Stephen Spottswood (Committee Member); Manoj Srinivasan (Committee Member)

Subjects:

Aerospace Engineering

Keywords:

Hypersonic; FTSI; fluid-thermal-structural interactions; aeroelasticity; aerothermoelasticity; time integration; loosely coupled

Witeof, ZacharyExploratory Study on the Design of Combined Aero-Thermo-Structural Experiments in High Speed Flows
Master of Science, The Ohio State University, 2013, Mechanical Engineering
This thesis describes potential configurations and responses of panels for experimental tests investigating fluid-thermal-structural interactions. The United States Department of Defense (DoD) and the National Aeronautics and Space Administration (NASA) are both interested in hypersonic vehicles due to their incredible capabilities for long range strike, surveillance and responsive access to space. These vehicles experience combined extreme aerodynamic heating and pressure loads. In addition, the structural loads are greatly affected by temperature-dependent material properties. Prediction of the response of these structures requires accurate, coupled fluid-thermal-structural analysis over long time records. High fidelity modeling over these long trajectories is prohibitively expensive computationally. In addition, validation of approximate models is limited due to the lack of experimental data capturing coupled fluid-thermal-structural interactions. The aim of this thesis is to explore the potential for experimental studies of fluid-thermal-structural interactions for validation and basic discovery purposes. Panel flutter is examined in this study, due to a simpler configuration for modeling. In order to accomplish the goal of this work, potential experimental facilities are explored to determine the conditions under which panels could be tested. An examination of common nondimensional parameters related to panel flutter is conducted to determine the utility of these parameters. A study of different panel configurations is conducted to determine the effect on panel response of varying flow conditions, boundary conditions and panel materials. Finally, the potential panel configurations that could be tested in two different facilities are discussed. It is found that there is a range of potential panel configurations that could be examined experimentally. For smaller facilities, the difference between the thinnest and thickest panels which exhibit a buckled response followed by the onset of instability is small and could lead to difficulties when manufacturing test articles. For larger facilities, this range is larger leading enabling more relaxed tolerances. In addition, it is found that while the structural boundary conditions have a significant effect on panel response, changing the thermal support conditions has a less significant effect on overall panel response. Varying thickness and panel material can result in larger thermal moments in the panel which can affect the utility of existing nondimensional parameters. It is also found that existing nondimensional parameters can provide some use in the sizing of panels for experimental tests. The overall conclusion is that limited experimental testing of fluid-thermal-structural interactions is feasible in existing facilities.

Committee:

Jack McNamara, PhD (Advisor); James Gregory, PhD (Committee Member)

Subjects:

Aerospace Engineering; Mechanical Engineering

Keywords:

Hypersonic; Fluid-Structure Interaction; Aerothermoelasticity; Panel Flutter; Fluid-Thermal-Structural Interaction

Sockalingam, SubramaniCoupling of Fluid Thermal Simulation for Nonablating Hypersonic Reentry Vehicles Using Commercial Codes FLUENT and LS-DYNA
MS, University of Cincinnati, 2008, Engineering : Mechanical Engineering
A frame work has been setup for the simulation of hypersonic reentry vehicles using commercial codes FLUENT and LS-DYNA. The main goal of this work was to set up a simple approach for the heat flux prediction and evaluation of the material thermal response during the reentry of the vehicle. Fluid thermal coupling for predicting the thermal response of a reusable non-ablating thermal protection system was set up. The computational fluid dynamics code (FLUENT) and the material thermal response codes (LS-DYNA) are loosely coupled to achieve the solution. The vehicle considered in the calculation is an axisymmetric vehicle flying at zero degree angle of attack. The frame work set up was validated with the results available in the literature. Good correlation was observed between the results from the commercial codes and the results from the literature. The mesh movement capability in LS-DYNA was implemented enabling future modeling of ablating thermal protection system.

Committee:

Ala Tabiei, PhD (Committee Chair); David Thompson, PhD (Committee Member); Prem Khosla, PhD (Committee Member); Kumar Vemaganti, PhD (Committee Member)

Subjects:

Mechanical Engineering

Keywords:

Hypersonic Reentry Vehicles; Thermal Protection System TPS; Ablation

Ling, YuNonlinear Response of a Skin Panel under Combined Thermal and Structural Loading
Master of Science, Miami University, 2012, Computational Science and Engineering

Future hypersonic vehicle will operate in an extreme environment, which involves extreme aerodynamic heating, fluctuating pressure and acoustic loading. Hypersonic vehicle must be reusable, lightweight and affordable in such environment. For hypersonic flight, the structure experiences complex aeroacoustic loads. The design of the structure depends on the ability to predict the response and the life of structure in extreme environment. This research presents a detailed investigation of the interactions and interplay among these parameters as evidenced by the nonlinear (static and dynamic) response of the panel.

A representative panel was selected as part of a ramp skin panel on a blend wing body hypersonic vehicle concept. This project focuses on the nonlinear response of the skin panel under combined thermal and structural loading. Thermal buckling, snap-through, and snap-buckling behaviors have been investigated by using different structural boundary condition. The long term goal of this research includes capturing fluid structure interaction as well as developing design curves for nonlinear response of the panel.

Committee:

Amit Shukla (Committee Chair); Edgar Caraballo (Committee Member); Kumar Singh (Committee Member); Timothy Cameron (Committee Member)

Subjects:

Mechanical Engineering

Keywords:

hypersonic aircraft; skin panel; nonlinear response; thermal buckling; snap-through

Culler, Adam JohnCoupled Fluid-Thermal-Structural Modeling and Analysis of Hypersonic Flight Vehicle Structures
Doctor of Philosophy, The Ohio State University, 2010, Aero/Astro Engineering

This dissertation describes coupled fluid-thermal-structural modeling and analysis of a semi-infinite insulated metallic panel and a blade-stiffened carbon-carbon skin panel for aerothermoelasticity and forced response prediction in hypersonic flow.

The United States Air Force' goals of affordable, reusable platforms capable of sustained hypersonic flight and responsive access to space depend on the ability to predict the response, the degradation, and ultimately the life of structures under combined, extreme aerodynamic heating and fluctuating pressure loads. However, the necessary modeling and prediction capabilities are severely limited in current commercial software due to the inability to seamlessly address multi-coupled, multi-scale fluid-thermal-structural interactions. Moreover, because of the complexity and expense of coupled computational methods, the capability is needed to define the necessary level of coupling a priori.

The aim of this dissertation is to develop coupled fluid-thermal-structural analysis methodology for response prediction in combined, extreme environments. Furthermore, it seeks to identify key characteristics (e.g., trajectories, operating conditions, and structural configurations) that determine the level of coupling needed for different situations. An additional focus is the targeted use of simplified temporal coupling strategies for reducing the computational expense of hypersonic aerothermoelasticity and forced response prediction over long durations.

First, in order to efficiently study the effects of fluid-thermal-structural coupling, an approximate hypersonic aerodynamics model is developed. The approximate model is verified and validated by comparison to aerodynamic pressure and heating data from hypersonic wind tunnel experiments and computational fluid dynamics solutions of the Navier-Stokes equations. Next, thermal and structural models of the panels are developed. The insulated metallic panel is modeled using the two-dimensional heat equation with a finite difference solution and von Karman plate theory with an assumed modes solution. Thermal and structural models of the carbon-carbon skin panel are developed using commercial finite element software.

Partitioned fluid-thermal-structural solution strategies are developed and used to systematically study the impact of multiple physical coupling mechanisms and simplified temporal coupling procedures. The aerothermoelastic behavior of the insulated metallic panel is found to be dependent on mutual coupling of aerodynamic heating and structural deformation. Additionally, it is determined that simplified temporal coupling procedures offer substantial reductions in computational expense, with negligible loss of accuracy, for aerothermoelastic analysis over long-duration hypersonic trajectories.

Quasi-static and transient dynamic structural responses of the carbon-carbon skin panel are investigated. It is found that the level of coupling needed for quasi-static response prediction depends largely on the in-plane structural boundary conditions, since greater resistance to thermal expansion results in larger deformations. Including these deformations in the aerodynamic heating analysis results in: nonuniform skin temperatures, asymmetric deformation, and elevated stresses. Predictions of panel failure and mode (static stress or snap-through) are found to be dependent on: trajectory, degree of coupling, and stiffness of in-plane boundary conditions. Additionally, it is determined that dynamic response predictions are sensitive to: mutual coupling of aerodynamic heating and structural deformation and temporal coupling of thermal and structural solutions. The degree of coupling needed for accurate dynamic response prediction increases with increasing fluctuating pressure levels and aerodynamic heating rates.

Committee:

Jack McNamara, PhD (Advisor); Jen-Ping Chen, PhD (Committee Member); Somnath Ghosh, PhD (Committee Member); Mo-How Shen, PhD (Committee Member)

Subjects:

Aerospace Materials; Design; Engineering; Mechanical Engineering

Keywords:

hypersonic; structures; multi-disciplinary; fluid-thermal-structural; coupled; thermal effects; aerothermoelasticity

Shakiba-Herfeh, MohammadModeling and Nonlinear Control of a 6-DOF Hypersonic Vehicle
Doctor of Philosophy, The Ohio State University, 2015, Electrical and Computer Engineering
In the past two decades, there has been a renewed and sustained effort devoted to modeling the dynamics of air-breathing hypersonic vehicles, for both simulation and control design purposes. The highly nonlinear characteristics of flight dynamics in hypersonic regimes and the consequent significance variability of the response with the operating conditions requires the development of innovative flight control solutions, hence the development of suitable model of the vehicle dynamics that are amenable to design, validation and rapid calibration of control algorithms. In this dissertation, a control-oriented and a simulation model of a generic hypersonic vehicle were derived to support the design and calibration of model-based flight controllers. A nonlinear robust adaptive controller was developed on the basis of the control-oriented model, that was shown to provide stable trajectory tracking in higher fidelity computer simulations. The first stage of this research was the development of a control design model (CDM) for the 6-degree-of-freedom dynamics of an air-breathing hypersonic aircraft based on an available high-fidelity first principle model. A method that incorporates the theory of compressible fluid dynamics and system identification methods, was proposed and implemented. The development of the CDM is based on curve fit approximation of the forces and moments acting on the vehicle, making the model suitable for control design. Kriging and Least Squares methods were used to find the most appropriate curve-fitted model of the aerodynamic forces for both the control design and the control simulation models. It was shown that the 6-DOF model can be both categorized as an under-actuated mechanical system, as well as an over-actuated system with respect to a chosen in- put/output pair of interest. An important contribution of this work is the development of a nonlinear adaptive controller for the 6-DOF control design model. The controller was endowed with a modular structure, comprised of an adaptive inner-loop attitude controller and a robust nonlinear outer-loop controller of fixed structure. The purpose of the outer- loop controller is to avoid the typical complexity of solutions derived from adaptive backstepping methods. A noticeable feature of the outer-loop controller is the presence of an internal model unit that generates the reference for the angle-of-attack, in spite of parametric model uncertainty. Airspeed, lateral velocity, vehicle’s heading and altitude were considered as regulated outputs of the system. Simulation results on the control simulation model show the effectiveness of the developed controller in spite of significant variation in the flight parameters.

Committee:

Andrea Serrani (Advisor); Vadim Utkin (Committee Member); Kevin Passino (Committee Member); Can Koksal (Committee Member)

Subjects:

Electrical Engineering

Keywords:

Nonlinear Control; Aerospace; 6DOF; Hypersonic; Control

Mirmasoudi, SaraHigh Temperature Transient Creep Analysis of Metals
Master of Science in Engineering (MSEgr), Wright State University, 2015, Mechanical Engineering
The ability to design vehicles capable of reaching hypersonic speeds has become a necessity to satisfy industry requirements, hence requiring the need for better understanding of creep behavior of materials. Although the steady state creep of metals has been analyzed rigorously, there is little known about transient creep of many metals. Understanding transient creep behavior of metals is crucial in analysis and design of short term hypersonic flight applications. Hence, a transient creep analysis of 304SS, Al7075-T6, Al2024-T6, Inconel 625, Inconel 718, and Rene N4 is carried out focusing on the microstructural behavior of these metals undergoing high temperature operating conditions. In doing so, the material properties that were unknown in literature were determined by parameter fitting techniques using existing steady state experimental data and also previous parametric studies determining critical parameters affecting strain values. A transient creep deformation map for each metal is produced including the required design space of the application.

Committee:

Mitch Wolff, Ph.D. (Committee Chair); Anthony Palazotto, Ph.D. (Advisor); Amir Farajian, Ph.D. (Committee Member)

Subjects:

Aerospace Engineering; Aerospace Materials; Engineering; Materials Science; Mechanical Engineering

Keywords:

Creep; steady state; transient; deformation; viscoplastic; hypersonic; short term; high temperature; design space; deformation mechanism maps;strain; stress; diffusion; power law creep; coefficient; Gibbs free energy; methodology; Inconel; AL; Rene N4

Clark, Daniel LeeLocally Optimized Covariance Kriging for Non-Stationary System Responses
Master of Science in Engineering (MSEgr), Wright State University, 2016, Mechanical Engineering
In this thesis, the Locally-Optimized Covariance (LOC) Kriging method is developed. This method represents a flexible surrogate modeling approach for approximating a non-stationary Kriging covariance structures for deterministic responses. The non-stationary covariance structure is approximated by aggregating multiple stationary localities. The aforementioned localities are determined to be statistically significant utilizing the Non-Stationary Identification Test. This methodology is applied to various demonstration problems including simple one and two-dimensional analytical cases, a deterministic fatigue and creep life model, and a five-dimensional fluid-structural interaction problem. The practical significance of LOC-Kriging is discussed in detail and is directly compared to stationary Kriging considering computational cost and accuracy.

Committee:

Ha-Rok Bae, Ph.D. (Advisor); Ramana Grandhi, Ph.D. (Committee Member); Joseph Slater, Ph.D., P.E. (Committee Member)

Subjects:

Applied Mathematics; Engineering; Mechanical Engineering

Keywords:

Kriging; surrogate modeling; optimization; fluid-structural interaction; FSI; Fatigue; Creep; Hypersonic; Ti-6242S; Analytical Life Model; Physics metamodeling; Non-stationary; Non-stationary Identification; localities; local surrogates

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