Formula 1 car front wings have evolved significantly over the last fifty years. Looking back at the past decade shows significant changes made due to rules and regulations by the Federation Internationale de l'Automobile and an increased understanding of aerodynamic concepts. There seems to be a trend where aerodynamic design concepts, previously seen in aviation, are being applied to Formula 1 front wings; this helps race teams increase downforce and reduce drag. This thesis analyzes these changes made over the past years and relates the material back to material that was learned by the aviation industry and attempts to synthesize conceptual Formula 1 front Wing designs using VORLAX, a vortex lattice panel method, used in the aviation industry. This insight would be beneficial for Formula 1 teams as there are budget and time restrictions applied to Computational Fluid Dynamic and wind tunnel testing, but panel methods are run in a matter of seconds as opposed to hours or days. So, if verified, preliminary designs can be rapidly tested to optimize the workflow and reduce the time required for Computational Fluid Dynamic and wind tunnel testing.

This work uses Arizona State University’s (ASU) newly developed high-speed vehicle stability and control screening methodologies to reverse-engineer famous United States Air Force (USAF) flight tests from the 1950s and 1960s. This thesis analyzes the root cause of Chuck Yeager's fateful 1953 supersonic spin in the Bell X-1A to become the "Fastest Man Alive". This thesis then takes a look back at Neil Armstrong's inadvertent atmospheric skip in the North American X-15 and his subsequent hypersonic flight months later. The fundamental flying qualities assessment shown in this work begins with calculating rigid-body frequencies and damping ratios of an aircraft to Military Standard (MIL) requirements, and uses these to create a full, classical stability and control analysis of a high-speed vehicle. Through reverse engineering the flight envelopes and missions for the above aircraft, it appears that the near-disasters of each flight were due to a confluence of then overlooked, yet fundamental, aerodynamic instabilities.

This thesis describes key Stability & Control parameters and a methodology for screening aircraft configurations for adequate handling qualities. For screening aircraft, a general-purpose Excel/ Visual Basic for Application (VBA) analysis tool was built. The analysis tool has built-in functionality to analyze aircraft utilizing a rudder for primary yaw control, collective horizontal tail for pitch, and either aileron or differential horizontal tail for roll control. Additionally, the tool transforms aerodynamic coefficients from the moment reference point to various center of gravity locations while saving data in Stability Axis and Body Axis; it also implements pitch trim. Key stability parameters of interest are the Short Period and Dutch Roll Frequencies, Roll and Spiral time constants, Cnβdynamic, Lateral Control Departure Parameter (LCDP), as well as the stick-fixed Short Period and Dutch Roll Damping. Other areas of interest include pitch and lateral-directional trim as well as the implementation of an Aileron-Rudder Interconnect system. This thesis uses the tool to analyze two historical: 1) the Bell X-2 and 2) the North American X-15 and two theoretical: 3) the “Sky Cruiser” and 4) the generic High-Speed Slender Aircraft (HSSA), aircraft. This thesis identifies varying Stability & Control problems between these aircraft and allows one to explore potential solutions to remedy their inherent flaws.

The Transonic Area Rule, developed by Richard T. Whitcomb in the early 1950s, revolutionized high-speed flight because its insight allowed engineers to reduce and/or delay the transonic drag rise. To this day, it is the rationale behind “coke-bottle” sculpturing (indenting the aircraft fuselage at the wing-fuselage junction) to alter the cross-sectional area development of the body. According to Whitcomb, this indentation is meant to create a smoother transition of cross-sectional area development of the body and consequently would reduce the number of shocks on the body, their intensity, and their shock pattern complexity. Along with this, modeling of a geometry’s transonic drag rise could be simplified by creating a comparable body of revolution with the same cross-sectional area development as the original geometry. Thus, the Transonic Area Rule has been advertised as an aerodynamic multitool. This new work probes the underlying mechanics of the Transonic Area Rule and determines just how accurate it is in producing its advertised results. To accomplish this, several different wave-drag approximation methods were used to replicate and compare the results presented in Whitcomb’s famous 1952 report16. These methods include EDET (Empirical Drag Estimation Technique)4, D2500 (Harris Wave Drag program)6, and CFD (Computational Fluid Dynamics) analysis through SU25. Overall drag increment data was collected for comparison with Whitcomb’s data. More in-depth analysis was then done on the flow conditions around the geometries using CFD solution plots. After analysis of the collected data was performed, it was discovered that this data argued against Whitcomb’s comparable body of revolution claim as no cases were demonstrated where the comparable body and original body yielded similar drag rise characteristics. Along with this, shock structures and patterns were not simplified in two of the three cases observed and were instead complicated even further. The only exception to this observation was the swept wing, cylindrical body in which all shocks were virtually eliminated at all observed Mach numbers. For the reduced transonic drag rise claim, the data argued in favor of this as the drag rise was indeed reduced for the three observed geometries, but only for a limited Mach number range.

The objective of this study is to estimate the variation of flight performance of a variable sweep wing geometry on the reverse engineered Boeing 2707-100 SST, when compared against the traditional delta wing approach used on supersonic airliner. The reason for this lies beneath the fact that supersonic orientations of wings doesn’t seem to work well for subsonic conditions, and subsonic wings are inefficient for supersonic flight. This would likely mean that flying long haul subsonic with supersonic wing geometry is inefficient compared to regular aircraft, but more importantly requires high takeoff/landing speeds and even long runways to bring the aircraft to hold. One might be able to get around this problem - partially - by adding thrust either by using afterburners, or by using variable geometry wings. To assess the flight performance, the research work done in this report focuses on implementing the latter solution to the abovementioned problem by using the aerodynamic performance parameters such as Coefficient of Lift, Coefficient of Drag along with its components specific to every test Mach number and altitude, along with the propulsion performance parameters such as thrust and thrust specific fuel consumption at different iterations of power settings of engine, flight Mach number and altitude in a propulsion database file to estimate flight performance using flight missions and energy-maneuverability theory approach. The flight performance was studied at several sweep angles of the aircraft to estimate the best possible sweep orientation based on the requirement of mission and an optimal flight mission was developed for an aircraft with swing wing capabilities.

This paper describes an effort to bring wing structural stiffness and aeroelastic considerations early in the conceptual design process with an automated tool. Stiffness and aeroelasticity can be well represented with a stochastic model during conceptual design because of the high level of uncertainty and variability in wing non-structural mass such as fuel loading and control surfaces. To accomplish this, an improvement is made to existing design tools utilizing rule based automated design to generate wing torque box geometry from a specific wing outer mold-line. Simple analysis on deflection and inferred stiffness shows how early conceptual design choices can strongly impact the stiffness of the structure. The impacts of design choices and how the buckling constraints drive structural weight in particular examples are discussed. The model is then carried further to include a finite element model (FEM) to analyze resulting mode shapes and frequencies for use in aeroelastic analysis. The natural frequencies of several selected wing torque boxes across a range of loading cases are compared.

Two methods of improving the life and efficiency of the Pulsed Inductive Thruster

(PIT) have been investigated. The first is a trade study of available switches to

determine the best device to implement in the PIT design. The second is the design

of a coil to improve coupling between the accelerator coil and the plasma. Experiments

were done with both permanent and electromagnets to investigate the feasibility of

implementing a modified Halbach array within the PIT to promote better plasma

coupling and decrease the unused space within the thruster. This array proved to

promote more complete coupling on the edges of the coil where it had been weak in

previous studies. Numerical analysis was done to predict the performance of a PIT

that utilized each suggested switch type. This model utilized the Alfven velocity to

determine the critical mass and energy of these theoretical thrusters.

The Doghouse Plot visually represents an aircraft’s performance during combined turn-climb maneuvers. The Doghouse Plot completely describes the turn-climb capability of an aircraft; a single plot demonstrates the relationship between climb performance, turn rate, turn radius, stall margin, and bank angle. Using NASA legacy codes, Empirical Drag Estimation Technique (EDET) and Numerical Propulsion System Simulation (NPSS), it is possible to reverse engineer sufficient basis data for commercial and military aircraft to construct Doghouse Plots. Engineers and operators can then use these to assess their aircraft’s full performance envelope. The insight gained from these plots can broaden the understanding of an aircraft’s performance and, in turn, broaden the operational scope of some aircraft that would otherwise be limited by the simplifications found in their Airplane Flight Manuals (AFM). More importantly, these plots can build on the current standards of obstacle avoidance and expose risks in operation.

This thesis explores the human factors effects pilots have when controlling the aircraft during the takeoff phase of flight. These variables come into play in the transitory phase from ground roll to flight, and in the initiation of procedures to abort a takeoff during the ground run. The FAA provides regulations for manufacturers and operators to follow, ensuring safe manufacture of aircraft and pilots that fly without endangering the passengers; however, details regarding accounting of piloting variability are lacking. Creation of a numerical simulation allowed for the controlled variation of isolated piloting procedures in order to evaluate effects on field performance. Reduced rotation rates and delayed reaction times were found to cause significant increases in field length requirements over values published in the AFM. A pilot survey was conducted to evaluate common practices for line pilots in the field, which revealed minimum regulatory compliance is exercised with little to no feedback on runway length requirements. Finally, observation of pilots training in a CRJ-200 FTD gathered extensive information on typical piloting timings in the cockpit. AEO and OEI takeoffs were observed, as well as RTOs. Pilots showed large variability in procedures and timings resulting in significant inconsistency in runway distances used as well as V-speed compliance. The observed effects from pilot timing latency correlated with the numerical simulation increased field length outputs. Variability in piloting procedures results in erratic field performance that deviates from AFM published values that invite disaster in an aircraft operating near its field performance limitations.

This thesis gives a detailed design process for a pulsed type thruster. The thrust stand designed in this paper is for a Pulsed Plasma Thruster built by Sun Devil Satellite Laboratory, a student organization at Arizona State University. The thrust stand uses a torsional beam rotating to record displacement. This information, along with impulse-momentum theorem is applied to find the impulse bit of the thruster, which varies largely from other designs which focus on using the natural dynamics their fixtures. The target impulse to record on this fixture was estimated to be 275 μN-s of impulse. Through calibration and experimentation, the fixture is capable of recording an impulse of 332 μN-s ± 14.81 μN-s, close to the target impulse. The error due to noise was characterized and evaluated to be under 5% which is deemed to be acceptable.