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Energy storage within hypersonic aircraft is becoming increasingly important with the development of more sophisticated electronic components and is an integral piece of expanding their overall capabilities. Hypersonics not only produce large external thermal loads, but also an abundance of internal thermal loads from components such as power electronics, avionics, and batteries. Additionally, limited volume within such vehicles introduces additional constraints. Thus, having efficient heat sinks that are capable of storing much of these heat loads is imperative. Passive thermal management systems, i.e., heat sinks, are preferable in most applications because they do not require power input to operate, and they are typically smaller than active systems such as coolant loops. In identifying and developing heat sinks with increased energy storage capability, an exhaustive search of available phase change materials (PCMs) is conducted. PCMs have been used in hypersonic vehicles in the past as a means of energy storage. Additionally, the use of energy-consuming endothermic reactions is considered. An innovative PCM-endothermic reaction hybrid approach is also developed. Both thermodynamic and transient/quasi-stationary models are developed for each of these proposed heat sink technologies. Prototypes are then developed for the best candidates to validate the models and draw conclusions on each heat sink's performance. Both the thermodynamic modeling and experimental results presented in this paper suggest that PCMs, endothermic reactions, and, especially, the hybrid system show greater energy storage capabilities than what is being used in hypersonic vehicles currently. Master of Science Hypersonic vehicles are an important topic of interest in the aerospace and defense industries. To be classified as hypersonic, a vehicle must travel at or above Mach 5, which is at least five times the speed of sound. Hypersonic vehicles often travel at high altitudes and a common application of the technology is in missiles. One major hurdle in developing hypersonic technologies at lower altitudes is that because of the high speeds, the outside skin temperature of the vehicle can reach thousands of degrees. Clearly, these temperatures can affect the heat load on the inside of the vehicle as can the thermal energy release of internal components such as the power electronics, the avionics, etc. To deal with these internal heat loads, innovative energy storage solutions are needed to efficiently and effectively store the thermal energy released internally. One approach considered here is the use of phase change materials (PCMs) as a storage medium. Melting such a material requires large amounts of energy and occurs at constant temperature. This is much more advantageous than heating a material in which only the temperature rises. Another approach considered in this thesis is that of using a chemical reaction, which requires energy input to proceed. Such a reaction is called an endothermic reaction and often results in a temperature decrease. Thus, simply mixing a set of reactants and adding energy helps cool the system. A final approach considered is a hybrid one, which combines a PCM material and an endothermic reaction. Such an approach combines the advantages of both. Each of these approaches are modeled thermodynamically to better understand how devices based on them work. Physical prototypes are then designed, built, and tested to confirm their performance. Both the modeling and experimental results presented in this thesis suggest that these devices show significantly improved energy storage capabilities over the devices currently used in hypersonic vehicles.

As the world moves towards environmental sustainability, the civil aviation enterprise has responded by setting challenging goals for significantly increased energy efficiency and reduced harmful emissions into the atmosphere as codified by National Aeronautics and Space Administration (NASA) and Advisory Council for Aircraft Innovation and Research in Europe (ACARE). The airline industry supports these goals because of their positive impact on operational cost and the environment. Achieving such goals requires introduction of novel technologies and aircraft concepts. Previous studies have shown that electrified aircraft can be effective in meeting these challenges.While there are several mechanisms to incorporate novel technologies for electrified aircraft, two such technologies: turbo-electric propulsion and distributed propulsion, are used in this research. Integration of these two technologies with the airframe leverages the well-known favorable interference between the wing and the tractor propeller wake to provide increased lift during takeoff.In the present research, the advantages and disadvantages of integrating a turbo-electric distributed propulsion (TEDiP) system are assessed for a regional transport aircraft (RTA). With near term motor technology, an improvement in trip fuel burn was observed on the four and six propeller variants of the TEDiP aircraft. The takeoff field length(TOFL) also improved in all three design variants which is a direct result of the working of distributed propulsion leading to better aerodynamic performance at takeoff conditions.The approach and findings for this research are reported in this thesis. Master of Science While air transportation system is considered the fastest means to travel, the avi-ation industry is responsible for 2.1% of all human-induced CO2 emissions, whichputs a renewed emphasis on environmental sustainability. There is heightenedinterest in exploring alternative propulsion technologies for aviation to mitigatethe effects of ever increasing demand for air travel coupled with fossil fuel pricevolatility.Ambitious plans have been outlined by leading aerospace organizations to reduceharmful emissions into the atmosphere. Achieving these ambitious goals requiresdevelopment and introduction of game changing technologies and aircraft con-cepts. Few such concepts include novel propulsion systems like all electric andhybrid-electric propulsion, distributed propulsion, and boundary layer ingestion.The X-57 is a novel all-electric aircraft being developed by NASA as a technologydemonstrator and makes use of multiple electric motors and propellers placedon the wing.Owing to battery technology limitations, all-electric and hybrid-electric propul-sion are not considered as viable options. In the near term, incorporatingdistributed propulsion alongside turbo-electric propulsion, for a Turbo-ElectricDistributed Propulsion(TEDiP) system may be a promising option in the near--to-mid-term. The overall goal of the present study is to investigate potentialbenefits and penalties of TEDiP systems for regional transport aircraft (RTA).To perform this study, the aerodynamics module of Pacelab Aircraft PreliminaryDesign (APD) Multi-Disciplinary Optimization (MDAO) framework is alteredto account for changes in wing-propeller aerodynamics due to the interactionof wing and multiple propellers. This required selection of a cost-effective toolthat captures aerodynamic data for multiple propellers and wing. VSPAEROis the aerodynamic tool of choice for this research. Aerodynamic data fromVSPAERO is coupled to APD and three TEDiP design variants with four, sixand eight propeller are designed with the ATR 72-500 as the baseline. Thebenefits and penalties of integrating the TEDiP system onto these variants isinvestigatedThe results show that a performance comparable to the baseline can be achievedin the near term with the four propeller variant even with current electricalsystems technology trends with a small weight penalty, and in the medium termon a six propeller variant. A decrease in trip fuel burn and improved takeofffield length(TOFL) performance justifies the usage of TEDiP systems.

Unmanned aircraft systems (UAS) often encounter turbulent fields that perturb the aircraft from its desired target trajectory, or in a manner that increases the load factor. The aircraft's fixed dihedral angle, providing passive roll-stiffness, is often selected based on lateral-directional stability requirements for the vehicle. A study to predict the effect of an active dihedral system on lateral-directional stability and vertical gust rejection capability was conducted to assess the performance and feasibility of the system. Traditionally, the dihedral location begins at the root to maintain wing structural requirements, however, the active dihedral system was also evaluated for dynamic stability and gust rejection performance at alternative dihedral breakpoint locations. Simulations were completed using linear parameter-varying (LPV) models, derived from traditional Newtonian aircraft dynamics and associated kinematic equations, to improve the modeling of the nonlinear active dihedral system. The stability of the LPV system was evaluated using Lyapunov stability theory applied to switched linear systems, assessing bounds of operation for the dihedral angle and flapping rate. An ideal feedback controller was developed using a linear–quadratic regulator (LQR) for both a discrete gust model and a continuous gust model, and a gain scheduled LQR controller was implemented to show the benefits of gain scheduling with a parameter varying state and input model. Finally, a cost analysis was conducted to investigate the real-world benefit of altering the dihedral breakpoint location. The effects of the active dihedral system on battery capacity and consumption efficiency were observed and compared with the gust rejection authority. Master of Science Unmanned aircraft systems (UAS) often encounter wind disturbances that perturb the aircraft from its desired target trajectory, or in a manner that increases the force encountered on the vehicle. The aircraft's fixed dihedral angle, providing stiffness to roll rotations, is often selected based on stability and control requirements for the vehicle. A study to predict the effect of a flapping wing (active dihedral) system on the stability, control, and wind gust rejection capability is completed to assess the performance and feasibility of such a system. Traditionally, the dihedral location begins at the root to maintain wing structural requirements, however, the active dihedral system was also evaluated for stability and wind gust rejection performance at alternative locations along the wing where the dihedral could begin, with intention of finding the best location. Simulations were completed using a varying set of simplified models, obtained from traditional aircraft mechanics, to improve the modeling of the true complex active dihedral system. The stability of the system was evaluated using various theories applied to the linear systems in attempt to define a bounded operating region for the dihedral angle and flapping motion. An ideal controller for the system was developed using ideas from well documented linear control theory for both a single wind gust and a continuous wind gust model. A controller that varies with vehicle flapping motion was implemented to show the benefits of scheduling the controller with a parameter varying state and input model. Finally, a cost analysis was conducted to investigate the real-world benefit of altering the dihedral starting location. The effects of the active dihedral system on battery capacity and consumption efficiency were observed and compared with the total gust rejection capability.

As the necessity of sustainable mobility rises, the demand to reduce the environmental impact of transporting mediums increases. The SUGAR Truss-Braced Wing (TBW) aircraft is a venture of Boeing, NASA and Virginia Tech for the N+3 generation of aircraft. These high-aspect-ratio aircraft are being designed with the aim to improve the structural and aerodynamic performance by implementing advanced technologies. Aerodynamics is a major factor influencing the performance of the aircraft, affecting the fuel consumption and emissions, especially due to drag. The multidisciplinary design optimization architecture for truss-braced-wing aircraft is dedicated to generate configurations with low fuel burn, maximum weight carrying capabilities and aircraft stability for long and medium range missions. The incorporation of flight dynamics at the conceptual design stage offers enhanced aerodynamic performance and wing flexibility for the aircraft. A robust flight dynamic system would need a detailed aerodynamic analysis of the aircraft with the focus on aeroelasticity. In this thesis, various aerodynamic coefficients and stability derivatives are investigated by applying Vortex-Lattice Method using OpenVSP, an open-source platform. The variation in aerodynamic parameters with changes in configurations and flow conditions are discussed as well. OpenVSP allows for study of these results with low computational expense. This will aid in efficient aerodynamic design and lay basis for flight dynamics analysis and its inclusion in the Multidisciplinary Design Analysis and Optimization (MDAO) framework. Master of Science The demand for sustainable mobility and green transportation is increasing. Reduction in the environmental impact of these mediums is the prime motivation for various research studies conducted in this domain. The SUGAR Truss-Braced Wing (TBW) aircraft configuration research, led by Boeing, NASA and Virginia Tech over the last two decades, aims at developing highly fuel-efficient next-generation aircraft. These high-aspect-ratio aircraft are being researched for improving the structural and aerodynamic performance by implementing advanced technologies. Aerodynamic performance of the aircraft influences the fuel consumption and emissions produced drastically. The current design optimization framework for the TBW aircraft focuses on development of these aircraft configurations with the goal to limit fuel burn and maximize payload carrying capability. Flight dynamics analysis can be significant to improve and obtain optimal solutions from the design process. Incorporation of flight dynamics at the conceptual design stage offers enhanced aerodynamic performance and wing flexibility for the next generation aircraft. Therefore, a detailed aerodynamic analysis of the aircraft would be needed to establish a systematic flight dynamics module. This thesis presents a new approach for formulating and analysing the aerodynamic coefficients and stability derivatives by implementing Vortex-Lattice Method available in the open-source software. This will further allow for inclusion of flight dynamics study of the new configurations for long and medium range missions within the existing framework.

This dissertation presents the development and application of a naval ship distributed system architecture framework, Architecture Flow Optimization (AFO), Dynamic Architecture Flow Optimization (DAFO), and Energy Storage System (ESS) model in naval ship Concept and Requirements Exploration (CandRE). The particular objective of this dissertation is to determine and assess Energy Storage System (ESS) capacity, charging and discharging capabilities in a complex naval ship system of systems to minimize vulnerability and maximize recoverability and effectiveness. The architecture framework is implemented through integrated Ship Behavior Interaction Models (SBIMs) that include the following: Warfighting Model (WM), Ship Operational Model (OM), Capability Model (CM), and Dynamic Architecture Flow Optimization (DAFO). These models provide a critical interface between logical, physical, and operational architectures, quantifying warfighting and propulsion capabilities through system measures of performance at specific capability nodes. This decomposition greatly simplifies the Mission, Power, and Energy System (MPES) design process for use in CandRE. AFO and DAFO are network-based, linear programming optimization methods used to design and analyze MPESs at a sufficient level of detail to understand system energy flow, define MPES architecture and sizing, model operations, reduce system vulnerability and improve system effectiveness and recoverability with ESS capabilities. AFO incorporates system topologies, energy coefficient component models, preliminary arrangements, and (nominal and damaged) steady state scenarios to minimize the energy flow cost required to satisfy all operational scenario demands and constraints. The refined DAFO applies the same principles as AFO, but adds two more capabilities, Propulsion and ESS charging, and maximizes effectiveness at each scenario timestep. DAFO also integrates with a warfighting model, operational model, and capabilities model that quantify the performance of tasks enabled by capabilities through system measures of performance at specific capability nodes. This dissertation provides a description of the design tools developed to implement these processes and methods, including a ship synthesis model, hullform exploration, MPES explorations and objective attribute metrics for cost, effectiveness and risk, using design of experiments (DOEs) response surface models (RSMs) and Energy Storage System (ESS) applications. Doctor of Philosophy This dissertation presents the development and application of a naval ship distributed system architecture framework, Architecture Flow Optimization (AFO), Dynamic Architecture Flow Optimization (DAFO), and Energy Storage System (ESS) design in naval ship Concept and Requirements Exploration (CandRE). The particular objective of this dissertation is to determine and assess Energy Storage System (ESS) capacity, charging and discharging capabilities in a complex naval ship system of systems to minimize vulnerability and maximize recoverability and effectiveness. The architecture framework is implemented through integrated Ship Behavior Interaction Models (SBIMs) that include the following: Warfighting Model (WM), Ship Operational Model (OM), Capability Model (CM), and Dynamic Architecture Flow Optimization (DAFO). These models provide a critical interface between logical, physical, and operational architectures, quantifying warfighting and propulsion capabilities through system measures of performance at specific capability nodes. This decomposition greatly simplifies the Mission, Power, and Energy System (MPES) design process for use in CandRE. AFO and DAFO are network-based, linear programming optimization methods used to design and analyze MPESs at a sufficient level of detail to understand system energy flow, define MPES architecture and sizing, model operations, reduce system vulnerability and improve system effectiveness and recoverability with ESS capabilities. AFO incorporates system topologies, energy coefficient component models, preliminary arrangements, and (nominal and damaged) steady state scenarios to minimize the energy flow cost required to satisfy all operational scenario demands and constraints. DAFO applies the same principles as AFO, but adds two more capabilities, Propulsion and ESS charging, and maximizes effectiveness at each scenario timestep. DAFO also integrates with a warfighting model, operational model, and capabilities model that quantify the performance of tasks enabled by capabilities through system measures of performance at specific capability nodes. This dissertation provides an overview of the design tools developed to implement these process and methods, including a ship synthesis model, hullform exploration, MPES explorations and objective attribute metrics for cost, effectiveness and risk, using design of experiments (DOEs) response surface models (RSMs) and Energy Storage System (ESS) applications.

This dissertation describes the development and application of a naval ship distributed system architectural framework, Architecture Flow Optimization (AFO), and Dynamic Architecture Flow Optimization (DAFO) to naval ship Concept and Requirements Exploration (CandRE). The architectural framework decomposes naval ship distributed systems into physical, logical, and operational architectures representing the spatial, functional, and temporal relationships of distributed systems respectively. This decomposition greatly simplifies the Mission, Power, and Energy System (MPES) design process for use in CandRE. AFO and DAFO are a network-based linear programming optimization methods used to design and analyze MPES at a sufficient level of detail to understand system energy flow, define MPES architecture and sizing, model operations, reduce system vulnerability and improve system reliability. AFO incorporates system topologies, energy coefficient component models, preliminary arrangements, and (nominal and damaged) steady state scenarios to minimize the energy flow cost required to satisfy all operational scenario demands and constraints. DAFO applies the same principles as AFO and adds a second commodity, data flow. DAFO also integrates with a warfighting model, operational model, and capabilities model that quantify tasks and capabilities through system measures of performance at specific capability nodes. This enables the simulation of operational situations including MPES configuration and operation during CandRE. This dissertation provides an overview of design tools developed to implement this process and methods, including objective attribute metrics for cost, effectiveness and risk, ship synthesis model, hullform exploration and MPES explorations using design of experiments (DOEs) and response surface models. Doctor of Philosophy This dissertation describes the development and application of a warship system architectural framework, Architecture Flow Optimization (AFO), and Dynamic Architecture Flow Optimization (DAFO) to warship Concept and Requirements Exploration (CandRE). The architectural framework decomposes warship systems into physical, logical, and operational architectures representing the spatial, functional, and time-based relationships of systems respectively. This decomposition greatly simplifies the Mission, Power, and Energy System (MPES) design process for use in CandRE. AFO and DAFO are a network-based linear programming optimization methods used to design and analyze MPES at a sufficient level of detail to understand system energy usage, define MPES connections and sizing, model operations, reduce system vulnerability and improve system reliability. AFO incorporates system templates, simple physics and energy-based component models, preliminary arrangements, and simple undamaged/damaged scenarios to minimize the energy flow usage required to satisfy all operational scenario demands and constraints. DAFO applies the same principles and adds a second commodity, data flow representing system operation. DAFO also integrates with a warfighting model, operational model, and capabilities model that quantify tasks and capabilities through system measures of performance. This enables the simulation of operational situations including MPES configuration and operation during CandRE. This dissertation provides an overview of design tools developed to implement this process and methods, including optimization objective attribute metrics for cost, effectiveness and risk.

Today, we see a renewed interest in aircraft with multiple propellers. To support conceptual design of these vehicles, one of the major needs is a fast and accurate method for estimating wing aerodynamic characteristics in the presence of multiple propellers. For the method to be effective, it must be easy to use, have rapid turnaround time and should be able to capture major wing–propeller interaction effects with sufficient accuracy. This research is primarily motivated by the need to assess the effectiveness of computational fluid dynamics (CFD) for simulating aerodynamic characteristics of wings with multiple propellers. The scope of the present research is limited to investigating the interaction between a single tractor propeller and a wing. This research aims to compare computational results from a Reynolds-Averaged Navier-Stokes (RANS) method, StarCCM+, and a vortex lattice method (VLM), VSP Aero. Two configurations that are analysed are 1) WIPP Configuration (Workshop for Integrated Propeller Prediction) 2) APROPOS Configuration. For WIPP, computational results are compared with measured lift and drag data for several angles of attack and Mach numbers. StarCCM+ results of wake flow field are compared with WIPP's wake survey data. For APROPOS, computed data for lift-to-drag ratio of the wing are compared with test data for multiple vertical and spanwise locations of the propeller. The results of the simulations are used to assess the effectiveness of the two CFD methods used in this research. Master of Science Today, we see a renewed interest in aircraft with multiple propellers due to an increasing demand for vehicles which fly short distances at low altitudes, be it flying taxis, delivery drones or small passenger aircrafts. To support conceptual design of vehicles, one of the major needs is a fast and accurate method for estimating wing aerodynamic characteristics in the presence of multiple propellers. For the method to be effective, it must be easy to use, have rapid turnaround time and should be able to capture major wing–propeller inter- action effects with sufficient accuracy. This research is primarily motivated by the need to assess the effectiveness of computational fluid dynamics (CFD) for simulating aerodynamic characteristics of wings with multiple propellers. Then only can we can take full advantage of the capabilities of the CFD methods and support design of emerging propeller driven air vehicles with an appropriate level of confidence. This research aims to compare high level methods with increasingly complex geometries and realistic models of physics like Reynolds Averaged Navier Stokes (RANS) and low level methods that rely on simplified geometry and simplified physics models like Vortex Lattice Methods (VLM). We will analyse multiple configurations and validate them against experi- mental data and thus assessing the effectiveness of the CFD models. This research investigates two configurations, 1) WIPP configuration 2) APROPOS configuration, for which experimental data is available. The results of the simulations are used to assess the effectiveness of the two CFD methods used in this research.

This investigation sought first to determine the feasibility of generating a surrogate model of the interference drag between nacelles and wing-fuselage systems suitable for the inclusion in a multidisciplinary design optimization (MDO) framework. The target aircraft was a single-aisle, transonic aircraft with a freestream Mach number of 0.8 at 35,000 feet and a design lift coefficient of 0.5. Using an MDO framework is necessary for placing the nacelle because of the competing objectives of the disciplines involved in aircraft design including structures, acoustics, and aerodynamics. A secondary goal was to determine what tools are necessary for accurately capturing interference drag effects on the system. This research used both Euler computational fluid dynamics (CFD) with a coupled viscous drag estimation tool and Reynolds Averaged Navier-Stokes (RANS) CFD to estimate the system drag. The initial trade space exploration that varied the nacelle location across a baseline airframe configuration was completed with the Euler solver, and it showed that appreciable overlap between the wing and nacelle led to large increases in interference drag. A follow-on study was conducted with RANS CFD where the wing shape was tailored for each unique nacelle position. In comparing the results of the Euler and the RANS CFD, it was determined that RANS is required to accurately capture the flow features. Euler solvers can create artifacts due to the lack of viscous effects within the model. Wing tailoring is necessary because of the sensitivity of transonic flows to geometric changes and the addition of neighboring components, such as a nacelle. The research showed that for above and aft wing locations, a nacelle can overlap the trailing edge without incurring a drag penalty. Nacelles placed in the conventional location, forward and beneath the wing, displayed low interference drag effects, as the nacelle had a small and local impact on the wing's aerodynamics. Given the high cost of computing a RANS solution with wing tailoring, and the large design space for nacelle locations, building a surrogate model for interference drag was found to be prohibitive at this time. As the cost of computing and mesh generation decreases, collecting the data for building a surrogate model may become tractable. Doctor of Philosophy Engine placement on an aircraft is dependent on multiple disciplines. Engine placement affects the noise of the aircraft because the wing can shield or reflect the engine noise. Engine placement impacts the structural loads of an aircraft, with some positions requiring more reinforcement that adds to the cost and weight of the aircraft. Aerodynamically, the engine placement impacts the vehicle's drag. Taken together, the only means of trading the different disciplines' needs is through a multidisciplinary design optimization (MDO) framework. The challenge of MDO frameworks is that they require numerous solutions to effectively explore the trade space. Thus, MDO frameworks employ fast, low-order tools to compute hundreds or thousands of different combinations of features. A common approach to make running MDO analysis feasible is to develop surrogate models of the key considerations. Current aerodynamic drag build-ups for aircraft do not consider the interference drag associated with engine placement. The first goal of this research was to determine the feasibility of generating a surrogate model for inclusion in an MDO framework. In order to collect the data required for the surrogate, appropriate tools to capture the interference drag are required. Building a surrogate requires a large number of samples, thus the aerodynamic solver must be fast, robust, and accurate. An Euler (inviscid) computational fluid dynamics (CFD) was used do explore the engine placement design space to test the feasibility of building the surrogate model. The target aircraft was a single-aisle, transonic aircraft with a freestream Mach number of 0.8, flying at an altitude of 35,000 feet and a design lift coefficient of 0.5. The initial vehicle used a baseline wing, and the engine placement was varied across the wing span and fuselage. The results showed that the conventional location, where the engine is forward and beneath the wing, had the a modestly beneficial interference drag, though positions near the trailing edge and above the wing also showed neutral interference drag. In general, if the engine overlapped the wing, the interference drag increased dramatically. A follow-on study used Reynolds Averaged Navier-Stokes (RANS) CFD to investigate seven engine placements above and aft of the wing. Each of these positions had the wing tailored such that the wing performance would be typical of a good transonic wing. The results showed that with wing tailoring, a moderate amount of overlap between the wing and nacelle results in reduced or neutral interference drag. This is in contrast with the baseline wing results that showed moderate overlap led to large increases in interference drag. The results from this research suggest that building a surrogate model of interference drag for transonic aircraft is not feasible given today's computational resources. In order to accurately model the interference drag, one must use a RANS CFD solver and tailor the wing. These requirements increase the cost of evaluating an engine position such that collecting enough for a surrogate model is prohibitively expensive. As computational speeds increase, and the ability to automate CFD mesh generation becomes less time intensive, the feasibility may increase. Using an Euler solver is insufficient because of the lack of viscous effects in the flow. The lack of a boundary layer leads to artifacts appearing in the flow when the nacelle and wing are in close proximity.

A time spectral and coupled adjoint based sensitivity analysis of rotor blade is carried out in this study. The time spectral method is an efficient technique to solve unsteady periodic problems by transforming unsteady equation of motion to a steady state one. Due to the availability of the governing equations in the steady form, the steady form of the adjoint equations can be applied for the sensitivity analysis of the coupled fluid-structure system. An expensive computational time and memory requirement for the unsteady adjoint sensitivity analysis is thus avoided. A coupled analysis of fluid, structural, and flight dynamics is carried out through a CFD/CSD/CA coupling procedure that combines FSI analysis with enforced trim condition. Coupled sensitivity analysis results and their validations are presented and compared with aerodynamics only sensitivity analysis results. The fluid-structure coupled adjoint based sensitivity analysis will be applied to the shape optimization of a rotor blade in the future work. Minimization of required power is the objective of the optimization problem with constraints on thrust and drag of the rotor. The bump functions are considered as the design variables. Rotor blade shape changes are obtained by using the bump function on the surface of the airfoil sections along the span. Doctor of Philosophy The work in this dissertation is motivated by the reducing the computational cost at the early design stage with guaranteed accuracy. In the research, the author proposes that the goal can be achieve through coupled adjoint based sensitivity analysis using a fluid structure interaction in time spectral form. Adjoint based sensitivity analysis is very efficient for solving design problems with a large number of design variables. The time spectral approach is used to overcome inefficient calculation of rotor flows by expressing flow and structural state variables as Fourier series with small number of harmonics. The accuracy and the efficiency of flow solver are examined by simulating UH-60A forward flight condition. A significant reduction in the computational cost is achieved by its Fourier series form of the periodic time response and the assumption of periodic steady state. A good agreement between time accurate and time spectral analysis is noted for the high speed forward flight condition of UH-60A configuration. Prediction from both methods also agree quite well with the experimental data. The adjoint based sensitivity analysis results are compared with the finite difference sensitivity analysis results. Even with presence of small discrepancies, these two results show a good agreement to each other. Coupled sensitivity analysis includes not only the effect of fluid state changes but also the contribution of structural deformation.

Work in this dissertation is motivated by reducing the design cost at the early design stage while maintaining high design accuracy throughout all design stages. It presents four key design methods to improve the performance of Efficient Global Optimization for multidisciplinary problems. First, a fidelity-calibration method is developed and applied to lower-fidelity samples. Function values analyzed by lower fidelity analysis methods are updated to have equivalent accuracy to that of the highest fidelity samples, and these calibrated data sets are used to construct a variable-fidelity Kriging model. For the design of experiment (DOE), a dynamic sampling method is developed and includes filtering and infilling data based on mathematical criteria on the model accuracy. In the sample infilling process, multi-objective optimization for exploitation and exploration of design space is carried out. To indicate the fidelity of function analysis for additional samples in the variable-fidelity Kriging model, a dynamic fidelity indicator with the overlapping coefficient is proposed. For the multidisciplinary design problems, where multiple physics are tightly coupled with different coupling strengths, multi-response Kriging model is introduced and utilizes the method of iterative Maximum Likelihood Estimation (iMLE). Through the iMLE process, a large number of hyper-parameters in multi-response Kriging can be calculated with great accuracy and improved numerical stability. The optimization methods developed in the study are validated with analytic functions and showed considerable performance improvement. Consequentially, three practical design optimization problems of NACA0012 airfoil, Multi-element NLR 7301 airfoil, and all-moving-wingtip control surface of tailless aircraft are performed, respectively. The results are compared with those of existing methods, and it is concluded that these methods guarantee the equivalent design accuracy at computational cost reduced significantly. Doctor of Philosophy In recent years, as the cost of aircraft design is growing rapidly, and aviation industry is interested in saving time and cost for the design, an accurate design result during the early design stages is particularly important to reduce overall life cycle cost. The purpose of the work to reducing the design cost at the early design stage with design accuracy as high as that of the detailed design. The method of an efficient global optimization (EGO) with variable-fidelity analysis and multidisciplinary design is proposed. Using the variable-fidelity analysis for the function evaluation, high fidelity function evaluations can be replaced by low-fidelity analyses of equivalent accuracy, which leads to considerable cost reduction. As the aircraft system has sub-disciplines coupled by multiple physics, including aerodynamics, structures, and thermodynamics, the accuracy of an individual discipline affects that of all others, and thus the design accuracy during in the early design states. Four distinctive design methods are developed and implemented into the standard Efficient Global Optimization (EGO) framework: 1) the variable-fidelity analysis based on error approximation and calibration of low-fidelity samples, 2) dynamic sampling criteria for both filtering and infilling samples, 3) a dynamic fidelity indicator (DFI) for the selection of analysis fidelity for infilled samples, and 4) Multi-response Kriging model with an iterative Maximum Likelihood estimation (iMLE). The methods are validated with analytic functions, and the improvement in cost efficiency through the overall design process is observed, while maintaining the design accuracy, by a comparison with existing design methods. For the practical applications, the methods are applied to the design optimization of airfoil and complete aircraft configuration, respectively. The design results are compared with those by existing methods, and it is found the method results design results of accuracies equivalent to or higher than high-fidelity analysis-alone design at cost reduced by orders of magnitude.

Enhancing performance of power electronics converters has always been an interesting topic in the power electronics community. Over the years, researchers and engineers are developing new high performance component, novel converter topologies, smart control methods and optimal design procedures to improve the efficiency, power density, reliability and reducing the cost. Besides pursuing high performance, researchers and engineers are striving to modularize the power electronics converters, which provides redundancy, flexibility and standardization to the end users. The trend of modularization has been seen in photovoltaic inverters, telecommunication power supplies, and recently, HVDC applications. A systematic optimal design approach for modular power converters is developed in this dissertation. The converters are developed for aerospace applications where there are stringent requirement on converter form factor, loss dissipation, thermal management and electromagnetic interference (EMI) performance. This work proposed an optimal design approach to maximize the nominal power of the power converters considering all the constraints, which fully reveals the power processing potential. Specifically, this work studied three-phase active front-end converter, three-phase isolated ac/dc converter and inverter. The key models (with special attention paid to semiconductor switching loss model), detailed design procedures and key design considerations are elaborated. With the proposed design framework, influence of key design variables, e.g. converter topology, switching frequency, etc. is thoroughly studied. Besides optimal design procedure, control issues in paralleling modular converters are discussed. A master-slave control architecture is used. The slave controllers not only follow the command broadcasted by the master controller, but also synchronize the high frequency clock to the master controller. The control architecture eliminates the communication between the slave controllers but keeps paralleled modules well synchronized, enabling a fully modularized design. Furthermore, the implementation issues of modularity are discussed. Although modularizing converters under form factor constraints adds flexibility to the system, it limits the design space by forbidding oversized components. This work studies the influence of the form factor by exploring the maximal nominal power of a double-sized converter module and comparing it with that of two paralleled modules. The tradeoff between modularity and performance is revealed by this study. Another implementation issue is related to EMI. Scaling up system capacity by paralleling converter modules induces EMI issues in both signal level and system level. This work investigates the mechanisms and provides solutions to the EMI problems. Ph. D. As penetration of power electronics technologies in electric power delivery keeps increasing, performance of power electronics converters becomes a key factor in energy delivery efficacy and sustainability. Enhancing performance of power electronics converters reduces footprint, energy waste and delivery cost, and ultimately, promoting a sustainable energy use. Over the years, researchers and engineers are developing new technologies, including high performance component, novel converter topologies, smart control methods and optimal design procedures to improve the efficiency, power density, reliability and reducing the cost of power electronics converters. Besides pursuing high performance, researchers and engineers are striving to modularize the power electronics converters, enabling power electronics converters to be used in a “plug-and-play” fashion. Modularization provides redundancy, flexibility and standardization to the end users. The trend of modularization has been seen in applications that process electric power from several Watts to Megawatts. This dissertation discusses the design framework for incorporating modularization into existing converter design procedure, synergically achieving performance optimization and modularity. A systematic optimal design approach for modular power converters is developed in this dissertation. The converters are developed for aerospace applications where there is stringent v requirement on converter dimensions, loss dissipation, and thermal management. Besides, to ensure stable operation of the onboard power system, filters comprising of inductors and capacitors are necessary to reduce the electromagnetic interference (EMI). Owning to the considerable weight and size of the inductors and capacitors, filter design is one of the key component in converter design. This work proposed an optimal design approach that synergically optimizes performance and promotes modularity while complying with the entire aerospace requirement. Specifically, this work studied three-phase active front-end converter, three-phase isolated ac/dc converter and three-phase inverter. The key models, detailed design procedures and key design considerations are elaborated. Experimental results validate the design framework and key models, and demonstrates cutting-edge converter performance. To enable a fully modularized design, control of modular converters, with focus on synchronizing the modular converters, is discussed. This work proposed a communication structure that minimizes communication resources and achieves seamless synchronization among multiple modular converters that operate in parallel. The communication scheme is demonstrated by experiments. Besides, the implementation issues of modularity are discussed. Although modularizing converters under form factor constraints adds flexibility to the system, it limits the design space by forbidding oversized components. This work studies the impact of modularity by comparing performance of a double-sized converter module with two paralleled modules. The tradeoff between modularity and performance is revealed by this study.

In recent years, there has been a growing shift in the world towards sustainability. For civil aviation, this is reflected in the goals of several organizations including NASA and ACARE as significantly increased fuel efficiency along with reduced harmful emissions in the atmosphere. Achieving the goals necessitates the advent of novel and radical aircraft technologies, NASA's X-57, is one such concept using distributed electric propulsion (DEP) technology. Although practical implementation of DEP is achievable due to the scale invariance of highly efficient electric motors, the current battery technology restricts its adoption for commercial transport aircraft. A Hybrid Electric Distributed Propulsion (HEDiP) system offers a promising alternative to the all-electric system. It leverages the benefits of DEP when coupled with a hybrid electric system. One of the areas needing improvement in HEDiP aircraft design is the fast and accurate estimation of wing aerodynamic characteristics in the presence of multiple propellers. A VLM based estimation technique was developed to address this requirement. This research is primarily motivated by the need to have mature conceptual design methods for HEDiP aircraft. Therefore, the overall research objective is to develop an effective conceptual design capability based on a proven multidisciplinary design optimization (MDO) framework, and to demonstrate the resulting capability by applying it to the conceptual design of a regional transport aircraft (RTA) with HEDiP systems. Master of Science Recent years have seen a growing movement to steer the world towards sustainability. For civil aviation, this is reflected in the goals of key organizations, such as NASA and ACARE, to significantly improve fuel efficiency, reduce harmful emissions, and decrease direct heat release in the atmosphere. Achieving such goals requires novel technologies along with radical aircraft concepts driven by efficiency maximization as well as using energy sources other than fossil fuel. NASA’s all-electric X-57 is one such concept using the Distributed Electric Propulsion (DEP) technology with multiple electric motors and propellers placed on the wing. However, today’s all-electric aircraft suffer from the heavy weight penalty associated with batteries to power electric motors. In the near term, a Hybrid Electric Distributed Propulsion (HEDiP) system offers a promising alternative. HEDiP combines distributed propulsion (DiP) technology powered by a mix of two energy sources, battery and fossil fuel. The overall goal of the present study is to investigate potential benefits of HEDiP systems for the design of optimal regional transport aircraft (RTA). To perform this study, the aerodynamics module of the Pacelab Aircraft Preliminary Design (APD) software system was modified to account for changes in wing aerodynamics due to the interaction with multiple propellers. This required the development of the Wing Aerodynamic Simulation with Propeller Effects (WASPE) code. In addition, a Wing Propeller Configuration Optimization (WIPCO) code was developed to optimize the placement of propellers based on location, number, and direction of rotation. The updated APD was applied to develop the HERMiT 2E series of RTA. The results demonstrated the anticipated benefits of HEDiP technologies over conventional aircraft, and provided a better understanding of the sensitivity of RTA designs to battery technology and level of hybridization, i.e., power split between batteries and fossil fuels. The HERMiT 6E/I was then designed to quantify the benefits of HEDiP systems over a baseline Twin Otter aircraft. The results showed that a comparable performance could be obtained with more than 50% saving in mission energy costs for a small weight penalty. The HERMiT 6E/I also requires only about 38% of the mission fuel borne by the baseline. This means a correspondingly lower direct atmospheric heat release, reduction in carbon dioxide and NOx emissions along with reduced energy consumptions.

Morphing aircraft offer many benefits. However, the design of stiff yet flexible structures still provides many obstacles to fully exploring and realizing morphing structures. Due to this, many morphing challenges remain open. Topology optimization is a type of structural optimization that optimizes the material layout of a structure based on imposed boundary conditions and load paths. This type of optimization is promising for solving morphing design challenges but many of the optimized structures are not suited for traditional manufacturing and material arrangements. Multi-material additive manufacturing is an emerging technology that can produce a single structure with many different materials integrated in custom geometries. This could be the solution to realizing topology optimized structures. Despite the rich amount of current research in morphing aircraft, many challenges still remain open and topology of morphing structures could provide the solution to these morphing challenges. Master of Science

A new flying wing design has been developed at the Unmanned Systems Lab (USL) at Virginia Tech to serve delivery and remote sensing applications in the developing world. The fully autonomous unmanned aerial vehicle (UAV), named EcoSoar, was designed with the goal of creating a business opportunity for local entrepreneurs in low-resource communities. The system was developed in such a way that local fabrication, operation, and maintenance of the aircraft are all possible. In order to present a competitive financial model for sustained drone services, EcoSoar is made with reliable low-cost materials and electronics. This paper lays out the rapid prototyping and flight experiment efforts that went into polishing the design, test results from an EcoSoar centered drone workshop in Kasungu, Malawi, and finally a range optimization study with flight test validation. Master of Science

This thesis describes the development of a nonlinear flight dynamics model for a small, fixed-wing unmanned aerial vehicle (UAV). Models developed for UAVs can be used for many applications including risk analysis, controls system design and flight simulators. Several challenges exist for system identification of small, low-cost aircraft including an increased sensitivity to atmospheric disturbances and decreased data quality from a cost-appropriate instrumentation system. These challenges result in difficulties in development of the model structure and parameter estimation. The small size may also limit the scope of flight test experiments and the consequent information content of the data from which the model is developed. Methods are presented to improve the accuracy of system identification which include data selection, data conditioning, incorporation of information from computational aerodynamics and synthesis of information from different flight test maneuvers. The final parameter estimation and uncertainty analysis was developed from the time domain formulation of the output-error method using the fully nonlinear aircraft equations of motion and a nonlinear aerodynamic model structure. The methods discussed increased the accuracy of parameter estimates and lowered the uncertainty in estimates compared to standard procedures for parameter estimation from flight test data. The significant contributions of this thesis are a detailed explanation of the entire system identification process tailored to the needs of a small UAV and incorporation of unique procedures to enhance identification results. This work may be used as a guide and list of recommendations for future system identification efforts of small, low-cost, minimally instrumented, fixed-wing UAVs. MS

Small Unmanned Aerial Vehicles (SUAVs) are rapidly being adopted in the National Airspace (NAS) but experience a much higher failure rate than traditional aircraft. These SUAVs are quickly becoming complex enough to investigate alternative methods of failure analysis. This thesis proposes a method of expanding on the Fault Tree Analysis (FTA) method to a Bayesian Belief Network (BBN) model. FTA is demonstrated to be a special case of BBN and BBN can allow for more complex interactions between nodes than is allowed by FTA. A model can be investigated to determine the components to which failure is most sensitive and allow for redundancies or mitigations against those failures. The introduced method is then applied to the Virginia Tech ESPAARO SUAV. Master of Science

Design and development of future high speed aircraft require the use of advanced modeling tools early on in the design phase to study and analyze complex aeroelastic, thermoelastic, and aerothermal interactions. This phase, commonly referred to as the conceptual design phase, involves using first principle based analytical models to obtain a practical starting point for the preliminary and detailed design phases. These analytical models are expected to, firstly, capture the effect of complex interactions between various subsystems using basic physics, and secondly, minimize computational costs. The size of a typical air-breathing hypersonic vehicle can vary anywhere between 12 ft, like the NASA X-43A, to 100 ft, like the NASP demonstrator vehicle. On the other hand, the performance expectations can vary anywhere between cruising at Mach 5 @ 85; 000 ft to Mach 10 @ 110; 000 ft. Reduction of computational costs is essential to efficiently sort through such a vast design space, while capturing the various complex interactions between subsystems has shown to improve accuracy of the design estimates. This motivates the need to develop modelling tools using first principle based analytical models with "needed" fidelity, where fidelity refers to the extent of interactions captured. With the advent of multidisciplinary design optimization tools, the need for an integrated modelling and analysis environment for high speed aircraft has increased substantially over the past two decades. The ever growing increase in performance expectations has made the traditional design approach of optimize first, integrate later obsolete. Designing a closed-loop control system for an aircraft might prove to be a difficult task with a geometry that yields an optimal (L/D) ratio, a structure with optimal material properties, and a propulsion system with maximum thrust-weight ratio. With all the subsystems already optimized, there is very little freedom for control designers to achieve their high performance goals. Integrated design methodologies focus on optimizing the overall design, as opposed to individual subsystems. Control-oriented modelling is an approach that involves making appropriate assumptions while modelling various subsystems in order to facilitate the inclusion of control design during the conceptual design phase. Due to their high lift-to-drag ratio and low operational costs, air-breathing hypersonic vehicles have spurred some interest in the field of high speed aircraft design over the last few decades. Modeling aeroelastic effects for such an aircraft is challenging due to its tightly integrated airframe and propulsion system that leads to significant deflections in the thrust vector caused by flexing of the airframe under extreme aerodynamic and thermal loads. These changes in the orientation of the thrust vector in turn introduce low frequency oscillations in the flight path angle, which make control system design a challenging task. Inclusion of such effects in the vehicle dynamics model to develop accurate control laws is an important part of control-oriented modeling. The air-breathing hypersonic vehicle considered here is assumed to be a thin-walled structure, where deformations due to axial, bending, shear, and torsion are modeled using the six independent displacements of a rigid cross section. Free vibration mode shapes are computed accurately using a novel scheme that uses estimates of natural frequency from the Ritz method as initial guesses to solve the governing equations using SUPORE, a two-point boundary value problem solver. A variational approach involving Hamilton's principle of least action is employed to derive the second order nonlinear equations of motion for the flexible aircraft. These nonlinear equations of motion are then linearized about a given cruise condition, modal analysis carried out on the linearized system, and the coupling between various significant modes studied. Further, open-loop stability analysis in time domain is conducted. Ph. D.

Multidisciplinary design optimization (MDO) requires many designs to be evaluated while searching for an optimum. As a result, the calculations done to evaluate the designs must be quick and simple to have a reasonable turn-around time. This makes aerodynamic calculations in the transonic regime difficult. Running computational fluid dynamics (CFD) calculations within the MDO code would be too computationally expensive. Instead, CFD is used outside the MDO to find two-dimensional aerodynamic properties of a chosen airfoil shape, BACJ, at a number of points over a range of thickness-to-chord ratios, free-stream Mach numbers, and lift coefficients. These points are used to generate surrogate models which can be used for the two-dimensional aerodynamic calculations required by the MDO computational design environment. Strip theory is used to relate these two-dimensional results to the three-dimensional wing. Models are developed for the center of pressure location, the lift curve slope, the wave drag, and the maximum allowable lift coefficient before buffet. These models have good agreement with the original CFD results for the airfoil. The models are integrated into the aerodynamic and aeroelastic sections of the MDO code. Master of Science

Thermal management has become a major concern in the design of current and future more and all electric aircraft (M/AEA). With ever increasing numbers of on-board heat sources, higher heat loads, limited and even decreasing numbers of heat sinks, integration of advanced intelligence, surveillance and reconnaissance (ISR) and directed energy weapons, requirements for survivability, the use of composite materials, etc., existing thermal management systems and their components have been pushed to the limit. To address this issue, more efficient methods of thermal management must be implemented to ensure that these new M/AEA aircraft do not overheat and prematurely abort their missions. Crucial to this effort is the need to consider advanced heat exchanger concepts, comparing their designs and performance with those of the conventional compact exchangers currently used on-board aircraft thermal management systems. As a step in this direction, the work presented in this thesis identifies two promising advanced heat exchanger concepts, namely, microchannel and phase change heat exchangers. Detailed conceptual design and performance models for these as well as for a conventional plate-fin compact heat exchanger are developed and their design and performance optimized relative to the criterion of minimum dry weight. Results for these optimizations are presented, comparisons made, conclusions drawn, and recommendations made for future research. These results and comparisons show potential performance benefits for aircraft thermal management incorporating microchannel and phase change heat exchangers. Master of Science

The primary objectives of this work are two-fold. First, additive manufacturing (AM) and rapid prototype (RP) testing are evaluated for use in production of a wind tunnel (WT) models. Second, an approach was developed to incorporate stability and control (SandC) WT data into aircraft conceptual design multidisciplinary design optimization (MDO). Both objectives are evaluated in terms of data quality, time, and cost. FDM(TM) and PolyJet AM processes were used for model production at low cost and time. Several models from a representative tailless configuration, ICE 101, were printed and evaluated for strength, cost and time of production. Furthermore, a NACA 0012 model with 20% chord flap was manufactured. Both models were tested in the Virginia Tech (VT) Open-Jet WT for force and moment acquisition. A 1/15th scale ICE 101 model was prepared for manufacturing, but limits of FDM(TM) technology were identified for production. An approach using WT data was adapted from traditional surrogate-based optimization (SBO), which uses computational fluid dynamics (CFD) for data generation. Split-plot experimental designs were developed for analysis of the WT SBO strategy using historical data and for WT testing of the NACA 0012. Limitations of the VT Open-Jet WT resulted in a process that was not fully effective for a MDO environment. However, resolution of ICE 101 AM challenges and higher quality data from a closed-section WT should result in a fully effective approach to incorporate AM and RP testing in an aircraft conceptual design MDO. Master of Science

The purpose of this study was to extend the analysis methods in PACELAB APD 3.1, a recent commercially available aircraft preliminary design tool with potential for MDO applications, for higher fidelity with physics-based instead of empirical methods and to enable the analysis of nonconventional aircraft configurations. The implementation of these methods was first validated against both existing models and wind tunnel data. Then, the original and extended PACELAB APD versions were used to perform minimum-fuel optimizations for both a traditional cantilever and strut-braced wing aircraft for a medium-range regional transport mission similar to that of a 737-type aircraft, with a minimum range of 3,115 nm and a cruise Mach number of 0.78. The aerodynamics, engine size / weight estimation and structural modules were heavily modified and extended to accomplish this. Comparisons to results for the same mission generated with FLOPS and VT MDO are also discussed. For the strut-braced configuration, large fuel savings on the order of 37% over the baseline 737-800 aircraft are predicted, while for the cantilever aircraft savings of 10-30% are predicted depending on whether the default or VT methods are utilized in the PACELAB analysis. This demonstrates the potential of the strut-braced configuration for reducing fuel costs, as well as the benefit of MDO in the aircraft conceptual design process. For the cantilever aircraft, FLOPS and VT MDO predict fuel savings of 8% and 23%, respectively. VT MDO predicts a fuel savings of 28% for the strut-braced aircraft over the baseline. Master of Science

This study explored a 130-passenger advanced turboprop commercial airliner with the purposes of economic feasibility and energy efficiency. A baseline vehicle and a derivative vehicle were researched and analyzed in detail. Based on the findings of this analysis, an advanced future airliner was designed. For the advanced airliner, advanced technologies were suggested and projections of these technology benefits were implemented. Detailed performance analysis was conducted for all three aircraft. The energy efficiency of each vehicle was compared to current and future N+3 aircraft. Lastly, cost analysis was performed to observe the impact of these energy savings. The three existing and future concepts evaluated were: 1) Bombardier 80- passenger Q400 baseline, 2) An expanded 130-passenger Bombardier Q400 termed the Q400XL, and 3) an N+3 advanced 130-passenger turboprop airliner termed the N+3 Airliner. The N+3 Airliner was compared to the SUGAR High, a Boeing/NASA N+3 aircraft, in both fuel efficiency and economic feasibility. The N+3 Airliner was 22 percent more energy efficient. At current oil prices, the N+3 Airliner had nearly identical operating cost. However, at two times current oil prices, the N+3 Airliner has a slight advantage economically. Therefore, as long as the price of oil is above 2011 oil prices, $3.03 per barrel, the N+3 Airliner will be an economically viable option. Ph. D.