Wireless sensor networks have become a rapidly growing research field in recent years. They are envisioned to have a wide range of applications in military, environmental and many other fields. We examine the performance of wireless sensor network applications to aeronautical telemetry. To date, test ranges have relied on a single telemetry ground station for the reception of packets from all air borne transmitters. We researched an alternate means of achieving this same goal with fewer resources. It is a well known fact that communication power and bandwidth are the most expensive commodities in wireless communications. The telemetry world is ever in need of ways and means to reduce power requirements of its networks while maximizing the use of available bandwidth. In our alternate method, packets will be reliably transported to a centrally located monitoring station in a series of hops. We will effectively reduce the power requirements of the network by minimizing the distance coverage of the sensor nodes. We will also explore different network topologies with a view to maximizing the use of available bandwidth. The alternate method will present a less expensive way to implement telemetry networks. Currently, telemetry networks make use of a single, huge and expensive base station receiving packets from all airplanes in test. Affordable sensor nodes placed strategically on the range and configured properly will achieve the same goal in a cost effective, power saving and bandwidth considerate manner.

Combustion emissions from aviation operations contribute significantly to climate change and air pollution. Accordingly, there is increasing interest in advancing battery-powered propulsion for aviation applications to reduce emissions. As batteries continue to improve, it is essential to recognize breakthroughs in battery specific energy in the context of air transport vehicles. Most electric aircraft designs and programs have focused on small aircraft because of restrictive battery performance. This work presents a feasibility assessment for an all-electric airliner based on an Airbus A220-100 with turbofan engines replaced by electric motors and propellers. The analysis compares the performance characteristics of the electric airliner to the A220-100 and establishes several configurations with varying battery pack-specific energy. The short-term electric airliner could replace conventional aircraft on very short, high-density missions. In contrast, the long-term electric airliner requires significant battery technology improvements that are not currently foreseen. The alternative long-term electric airliner could complete half of the A220-100’s missions, but the necessary specific energy value is also not anticipated shortly. All-electric airliners would significantly impact manufacturing, operations, costs, and emissions but are commercially infeasible with current battery technology. Additional development of more advanced battery technology is required to increase the specific energy of battery packs, enhance battery safety and reliability, and develop lighter high-power electric motors.

Investigation of complex multirotor aerodynamic phenomena via wind tunnel experimentation is becoming extremely important with the rapid progress in advanced distributed propulsion VTOL concepts. Much of this experimentation is being performed in large, highly advanced tunnels. However, the proliferation of this class of vehicles extends to small aircraft used by small businesses, universities, and hobbyists without ready access to this level of test facility. Therefore, there is a need to investigate whether multirotor vehicles can be adequately tested in smaller wind tunnel facilities. A test rig for a 2.82-pound quadcopter was developed to perform powered testing in the Cal Poly Aerospace Department’s Low Speed Wind Tunnel, equipped with a 3-foot tall by 4-foot wide test section. The results were compared to data from similar tests performed in the U.S. Army 7-by 10-ft Wind Tunnel at NASA Ames. The two data sets did not show close agreement in absolute terms but demonstrated similar trends. Due to measurement uncertainties, the contribution of wind tunnel interference effects to this discrepancy in measurements was not able to be properly quantified, but is likely a major contributor. Flow visualization results demonstrated that tunnel interference effects can likely be minimized by testing at high tunnel speeds with the vehicle pitched 10-degrees or more downward. Suggestions towards avoiding the pitfalls inherent to multirotor wind tunnel testing are provided. Additionally, a modified form of the conventional lift-to-drag ratio is presented as a metric of electric multirotor aerodynamic efficiency.

With the pursuit of more advanced and environmentally-friendly technologies of today’s society, the airline industry has been pushed further to investigate solutions that will reduce airport noise and congestion, cut down on emissions, and improve the overall performance of aircraft. These items directly influence airport size (runway length), flight patterns in the community surrounding the airport, cruise speed, and many other aircraft design considerations which are setting the requirements for next generation aircraft. Leading the research in this movement is NASA, which has set specific goals for the next generation regional airliners and has categorized the designs that meet the criteria as Cruise Efficient Short Takeoff and Land (CESTOL) aircraft. With circulation control (CC) technology addressing most of the next generation requirements listed above, it has recently been gaining more interest, thus the basis of this research. CC is an active flow control method that uses a thin sheet of high momentum jet flow ejected over a curved trailing edge surface and in turn utilizes Coanda effect to increase the airfoil’s circulation, augmenting lift, drag, and pitching moment. The technology has been around for more than 75 years, but is now gaining more momentum for further development due to its significant payoffs in both performance and system complexity. The goal of this research was to explore the design of the CC flap shape and how it influences the local flow field of the system, in attempt to improve the performance of existing CC flap configurations and provide insight into the aerodynamic characteristics of the geometric parameters that make up the CC flap. Multiple dual radius flaps and alternative flap geometry, prescribed radius, flaps were developed by varying specific flap parameters from a baseline dual radius flap configuration that had been previously developed and researched. The aerodynamics of the various flap geometries were analyzed at three different flight conditions using two-dimensional CFD. The flight conditions examined include two low airspeed cases with blown flaps at 60° and 90° of deflection, and a transonic cruise case with no blowing and 0° of flap deflection. Results showed that the shorter flaps of both flap configurations augmented greater lift for the low airspeed cases, with the dual radius flaps producing more lift than the corresponding length prescribed radius. The large lift generation of these flaps was accompanied by significant drag and negative pitching moments. The incremental lift per drag and moment produced was best achieved by the longer flap lengths, with the prescribed radius flaps out-performing each corresponding dual radius. Longer flap configurations also upheld the better cruise performance with the least amount of low airspeed flow, drag, and required angle of attack for a given cruise lift coefficient. The prescribed radius flaps also presented a favorable trait of keeping a more continuous skin friction distribution over the flap when the flaps were deflected, where all dual radius configurations experienced a distinct fluctuation at the location where the surface curvature changes between its two radii. The prescribed radius flaps displayed a similar behavior when the flaps were not deflected, during the cruise conditions analyzed. Performance trends for the different flap configurations, at all three flight conditions, are presented at the end of each respective section to provide guidance into the design of CC geometry. The results of the presented research show promise in modifying geometric surface parameters to yield improved aerodynamics and performance.

An automated computer simulation of the two-dimensional planar Cal Poly Supersonic Ejector test rig is developed. The purpose of the simulation is to identify the operating conditions which produce the saturated, Fabri choke and Fabri block aerodynamic flow patterns. The effect of primary to secondary stagnation pressure ratio on the efficiency of the ejector operation is measured using the entrainment ratio which is the secondary to primary mass flow ratio. The primary flow of the ejector is supersonic and the secondary (entrained) stream enters the ejector at various velocities at or below Mach 1. The primary and secondary streams are both composed of air. The primary plume boundary and properties are solved using the Method of Characteristics. The properties within the secondary stream are found using isentropic relations along with stagnation conditions and the shape of the primary plume. The solutions of the primary and secondary streams iterate on a pressure distribution of the secondary stream until a converged solution is attained. Viscous forces and thermo-chemical reactions are not considered. For the given geometry the saturated flow pattern is found to occur below stagnation pressure ratios of 74. The secondary flow of the ejector becomes blocked by the primary plume above pressure ratios of 230. The Fabri choke case exists between pressure ratios of 74 and 230, achieving optimal operation at the transition from saturated to Fabri choked flow, near the pressure ratio of 74. The case of optimal expansion yields an entrainment ratio of 0.17. The entrainment ratio results of the Cal Poly Supersonic Ejector simulation have an average error of 3.67% relative to experimental data. The accuracy of this inviscid simulation suggests ejector operation in this regime is governed by pressure gradient rather than viscous effects.

Multi-Body System Simulation combined with System Identification was developed as a method for determining the stability characteristics of a human powered helicopter(HPH) configurations. HPH stability remains a key component for meeting competition requirements, but has not been properly treated. Traditional helicopter dynamic analysis is not suited to the HPH due to its low rotation speeds and light weight. Multi-Body System Simulation is able to generate dynamic response data for any HPH configuration. System identification and linear stability theory are used to determine the stability characteristics from the dynamic response. This thesis focuses on the method development and doesn't present any HPH analysis results.

The ability to predict the distance required for an aircraft to takeoff is an essential component of aircraft design. It involves aspects related to each of the major aircraft systems: aerodynamics, propulsion, configuration, structures, and stability and control. For an aircraft designed for short takeoffs and landings (STOL), designing the aircraft to provide a short takeoff distance, or more precisely the balanced field length (BFL), often leads to the use of a powered lift technique such as circulation control (CC). Although CC has been around for many years, it has never been used on a production aircraft. This is in part due to the lack of knowledge as to how well CC can actually perform as a high lift device. This research provides a solution to this problem. By utilizing high fidelity computational fluid dynamics (CFD) aerodynamic data, a four-dimensional design space which was populated and modeled using a Monte Carlo approach, and a Gaussian Processes regression technique, an effective aerodynamic model for CC was produced which was then used in a BFL simulation. Three separate models were created of increasing quality which were then used in the BFL performance calculations. A comprehensive gridding methodology was provided as well as computational and grid dependence error analysis. Specific consideration was given to the effect of resolving the turbulent boundary layer in both the gridding and solving processes. Finally, additional turbulence model validation work was performed, both to match previously performed experimental data and to provide a comparison of different models’ abilities to predict separation.

Once a satellite has completed its operational period, it must be removed responsibly in order to reduce the risk of impacting other missions. Geostationary Transfer Orbits (GTOs) offer unique challenges when considering disposal of spacecraft, as high eccentricity and orbital energy give rise to unique challenges for spacecraft designers. By leveraging small satellite research and integration techniques, a deployable drag sail module was analyzed that can shorten the expected orbit time of launch vehicle stages in GTO. A tool was developed to efficiently model spacecraft trajectories over long periods of time, which allowed for analysis of an object’s expected lifetime after its operational period had concluded. Material limitations on drag sail sizing and performance were also analyzed in order to conclude whether or not a system with the required orbital performance is feasible. It was determined that the sail materials and configuration is capable of surviving the expected GTO environment, and that a 49 m2 drag sail is capable of sufficiently shortening the amount of time that the space vehicles will remain in space.

The Cal Poly SLO Turbo-Prop Racer project aims to design a world speed record-breaking aircraft, capable of flying more than 550 miles per hour on a 3-kilometer closed course. To further this endeavor, this thesis presents the calculations of load distributions across the aircraft’s wing and tail and preliminary structural estimates of primary structural components for verification of the loads calculations and for use in a future finite element model. The aircraft’s fundamental design characteristics effect on the structure of the aircraft, namely the unique Y-tail design, are first examined. Then, loads are calculated in accordance with the regulation dictated by CS-23. Maneuvering loads, gust loads, ground loads, and engine loads are calculated through the Vortex-Lattice Method and CS-23 to provide input for detailed structural analysis. Structural thickness estimates are found using simplified analytical stress analysis. The wing and tail’s primary spars’ spar-caps and shear-webs, the wing and tail skins, and the rear fuselage are all calculated. The loads and thicknesses found are shown to be within order of reason and to support the fundamental design characteristics of the aircraft, pushing the project to continue toward its goal.

The structural design of flight vehicles is a balancing act between maximizing loading capability while minimizing weight. An engineer must consider not only the classical static structural yielding failure of a vehicle, but a variety of ways in which structural deformations can in turn, affect the loading conditions driving those deformations. Lift redistribution, divergence, and flutter are exactly such dynamic aeroelastic phenomena that must be properly characterized during the design of a vehicle; to do otherwise is to risk catastrophe. Relevant within the university context is the design of small-scale aircraft for student projects and of particular consideration, the DBF competition hosted by AIAA. This work implements a variety of aeroelastic analysis methods: K and PK with Theodorsen aerodynamics via Matlab, NASA EZASE, and the FEMAP NX NASTRAN Aeroelasticity Package. These techniques are applied to a number of baseline test cases in addition to two representative DBF wings. Both wings considered ultimately indicated stability within reasonable flight conditions, although each for a different reason. Analysis results for the Cal Poly 2020 wing, a spar-rib construction emblematic of the collocation design approach, showed that the wing was stable within expected flight regions. The USC 2020 wing model, a composite top spar construction, exhibited unstable behavior, however this was well outside the scope of expected flight conditions. The codebase developed as a part of this work will serve as a foundation for future student teams to perform aeroelastic analyses of their own and support continued aeroelastic research at Cal Poly - SLO.