Abstract The main objective of this thesis was to develop a method than can be used to approximate the pressure forces on air vehicles traveling at hypersonic speed (Mach number > 5). The aerodynamic forces such as lift and drag were calculated from the pressure values on the surface of the airplane. Pitching moment was also tabulated. This work was initiated based on the idea of developing a flow solver proficient and capable of providing aerodynamic data (lift and drag look-up tables) for hypersonic air vehicles that can be fed to a flight simulator (used by the Aviation Systems Department) at the University of Tennessee Space Institute. Several approximation methods are used to solve hypersonic such as shock expansion method. Based on different studies, Computational Fluid Dynamic (CFD) proved to produce very accurate results; however, it is a difficult technique to use. In this thesis work Newtonian Method was adopted as a technique to approximate the aerodynamic forces and hence the performance of hypersonic airplanes, therefore, a computer program (Hyper-N) has been developed for aerodynamic analysis of three dimensional geometries airplane. The program is designed to read in a previously configured list of plates and compute the aerodynamic forces and moments for hypersonic free stream conditions. Programming was completed using MatLab language. The results obtained from the Hyper-N program were for the experimental airplane X-43A which were found to match the results when the shock expansion method is used for the same airplane, [1]. Because of the difficulties involve in using CFD or the complete Navier Stocks equation to obtain the aerodynamic forces on bodies traveling at hypersonic speeds, the Newtonian method is considered to be the most efficient technique to use for preliminary evaluation of the performance of hypersonic airplanes. Modified Newtonian theory and the computational requirement of the code are described. A number of geometric configurations, including the X-43A (experimental hypersonic) airplane, are provided as examples of applications of the Hyper-N program.

A flight data acquisition system was developed to aid unmanned vehicle designers in verifying the vehicle's design performance. The system is reconfigurable and allows the designer to choose the correct combination of complexity, risk, and cost for a given flight test. The designer can also reconfigure the system to meet packaging and integration requirements. System functionality, repeatbility, and accuracy was validated by collecting data during multiple flights of a radio-controlled aircraft. Future work includes sensor fusion, thrust prediction methods, stability and control derivative estimation, and growing Cal Poly's small-scale component aerodynamic database.

In order to accurately conduct the preliminary and detailed design of solar powered Unmanned Aerial Systems (UAS), it is necessary to have a thorough understanding of the systems involved. In particular, it is desirable to have mathematical models and analysis tools describing the energy income and expenditure of the vehicle. Solar energy income models may include available solar irradiance, photovoltaic array power output, and maximum power point tracker efficiency. Energy expenditure models include battery charging and discharging characteristics, propulsion system efficiency, and aerodynamic efficiency. In this thesis, a series of mathematical models were developed that characterize the performance of these systems. Several of these models were then validated against test data. Testing was conducted on specific components used by a solar UAS designed and built by students at the California Polytechnic State University, San Luis Obispo, which completed a six-hour flight relying only on solar energy in May 2019. Results indicate that, while some models accurately predicted test outcomes, others still need further improvement. While these models may be useful during the preliminary and detailed design phases of a solar powered UAS, specific component testing should be conducted to converge on the most desired design solution.

A method for sizing remotely powered unmanned aerial vehicles is presented to augment the conventional design process. This method allows for unconventionally powered aircraft to become options in trade studies during the initial design phase. A design matrix is created that shows where, and if, a remotely powered vehicle fits within the design space. For given range and power requirements, the design matrix uses historical data to determine whether an internal combustion or electrical system would be most appropriate. Trends in the historical data show that the break in the design space between the two systems is around 30 miles and 1 kW. Electrical systems are broken into subcategories of onboard energy sources and remote power sources. For this work, only batteries were considered as an onboard energy source, but both lasers and microwaves were considered for remote power transmission methods. The conventional sizing method is adjusted to so that it is based on energy consumption, instead of fuel consumption. Using the manner in which microwaves and laser propagate through the atmosphere, the weight fraction of a receiving apparatus is estimated. This is then used with the sizing method to determine the gross takeoff weight of the vehicle. This new sizing method is used to compare battery systems, microwave systems, and laser systems.

The purpose of this work is to help bridge the gap between aircraft conceptual design and analysis. Much work is needed, but distilling essential characteristics from a design and collecting them in an easily accessible format that is amenable to use by inexpensive analysis tools is a significant contribution to this goal. Toward that end, four types of reduced-fidelity or degenerate geometric representations have been defined and implemented in VSP, a parametric geometry modeler. The four types are degenerate surface, degenerate plate, degenerate stick, and degenerate point, corresponding to three-, two-, one-, and zero- dimensional representations of underlying geometry, respectively. The information contained in these representations was targeted specifically at lifting line, vortex lattice, equivalent beam, and equivalent plate theories, with the idea that suitability for interface with these methods would imply suitability for use with many other analysis techniques. The ability to output this information in two plain text formats— comma separated value and Matlab script—has also been implemented in VSP, making it readily available for use. A modified Cessna 182 wing created in VSP was used to test the suitability of degenerate geometry to interface with the four target analysis techniques. All four test cases were easily completed using the information contained in the degenerate geometric types, and similar techniques utilizing different degenerate geometries produced similar results. The following work outlines the theoretical underpinnings of degenerate geometry and the fidelity-reduction process. It also describes in detail how the routines that create degenerate geometry were implemented in VSP and concludes with the analysis test cases, stating their results and comparing results among different techniques.

Aerodynamic topology optimization is a useful tool in the aerodynamic design pro-cess, especially when looking for marginal gains within a design. One example isa turboprop racer concept aircraft that is designed with the goal of breaking worldspeed records. An optimization framework was developed with the intention of laterbeing applied to this design. In the early design stages, the optimization frameworkmust focus on quicker methods of drag estimation, such as a panel codes. The largenumber of design variables in topology optimization can exponentially increase func-tion evaluations and thus computational cost. A vorticity-based panel solver wasproven out for this application to reduce the computational cost while keeping theaccuracy of the results similar to that of traditional CFD solvers in conditions with-out prominent flow separation. The framework developed here includes geometryparameterization, function evaluation scripting, and post-processing, which are allrun within the optimization algorithm. The designs used to validate the solver arewing-fuselage systems of various sailplane configurations with existing experimentaldata. These sailplane designs were also used as the initial geometry to demonstratethe framework. A parametric optimum was found to reduce drag by 9%, but it mustbe noted that this method does have certain trade offs and limitations.

The aim of this effort is to implement a nonlinear flight control architecture, specifically flight path control via command filtered backstepping, for use in AME UAS's Fury® 1500 unmanned flying wing aircraft. Backstepping is a recursive, control-effort minimizing, constructive design procedure that interlaces the choice of a Lyapunov function with the design of feedback control. It allows the use of certain plant states to act as intermediate, virtual controls, for others breaking complex high order systems into a sequence of simpler lower-order design tasks. Work herein is a simplified implementation based on publications by Farrell, Sharma, and Polycarpou. Online approximation is not applied, however command filtering along with two variants of control allocation is. This minimal approach was done to mitigate risk, as adaptation could be added in future work to this baseline. Command filtering assures that control inputs generated meet magnitude, rate, and bandwidth constraints for states and actuators as well as provides command derivatives that reduce computation. Two different forms of control allocation were implemented, the simplest a least-squares pseudo-inverse and the second an optimal quadratic programming method. Two Simulink based simulations successfully flew AME's Fury® 1500 UAS: a nominal case with fully operational actuators and a failure case with an actuator stuck at -10°. Coordinated flight for both cases with outer loop tracking was achieved for a demanding autopilot task of simultaneously varying heading and flight-path angle commands, ±60° and ±10° respectively, for a constant airspeed command of 135 ft/s. Command signals were generated were achievable due to the command filter implementation.

Various methods have been used, including airborne radars, LIDAR, observation of flying birds, towers, tethered balloons, and aircraft to gain both a qualitative and quantitative representation of how heat and moisture are transported to higher altitudes and grow the boundary or mixing layer by thermal updrafts. This paper builds upon that research using an instrumented glider to determine the structure and build a mathematical model of thermals in a desert environment. During these flights, it was discovered that the traditional view of a thermal as a singular rising plume of air did not sufficiently explain what was being observed, but rather another phenomenon was occurring. This paper puts forth the argument and a mathematical model to show that thermals actually take the form of a hexagonal convection cell at higher levels in the convective boundary layer when the thermal acts as if unrestrained by borders as in non-linear cases of free convection.

The purpose of this study is to investigate the execution and validity of various predictive methods used in the design of the aerodynamic pod housing NASA’s Marshall Airborne Polarimetric Imaging Radiometer (MAPIR) on the University of Tennessee Space Institute’s Piper Navajo research aircraft. Potential flow theory and wing theory are both used to analytically predict the lift the MAPIR Pod would generate during flight; skin friction theory, empirical data, and induced drag theory are utilized to analytically predict the pod’s drag. Furthermore, a simplified computational fluid dynamics (CFD) model was also created to approximate the aerodynamic forces acting on the pod. A limited flight test regime was executed to collect data on the actual aerodynamic effects of the MAPIR Pod. Comparison of the various aerodynamic predictions with the experimental results shows that the assumptions made for the analytic and CFD analyses are too simplistic; as a result, the predictions are not valid. These methods are not proven to be inherently flawed, however, and suggestions for future uses and improvements are thus offered.

Various methods have been used, including airborne radars, LIDAR, observation of flying birds, towers, tethered balloons, and aircraft to gain both a qualitative and quantitative representation of how heat and moisture are transported to higher altitudes and grow the boundary or mixing layer by thermal updrafts. This paper builds upon that research using an instrumented glider to determine the structure and build a mathematical model of thermals in a desert environment. During these flights, it was discovered that the traditional view of a thermal as a singular rising plume of air did not sufficiently explain what was being observed, but rather another phenomenon was occurring. This paper puts forth the argument and a mathematical model to show that thermals actually take the form of a hexagonal convection cell at higher levels in the convective boundary layer when the thermal acts as if unrestrained by borders as in non-linear cases of free convection.