<div><div><div><p>The advent of nano and micro class satellites has generated new demand for compact and efficient propulsion systems. Traditional propulsion technologies have been miniaturized for the CubeSat platform and new technology solutions have been proposed to address this demand. However, each of these approaches has disadvantages when applied within the context of a CubeSat. One potential low mass and power alternative is Film-Evaporation MEMS Tunable Array (FEMTA) micropropulsion which is capable of generating 150μN of thrust using 0.65W of electrical power and ultra-pure deionized water as propellant. The FEMTA thruster is etched into a 1cm × 1cm × 0.3mm silicon substrate using standard photolithography and microfabrication techniques. Each thruster consists of a 4 μm wide nozzle and platinum resistive heaters. Capillary pressure prevents the water from leaking through the nozzle and the heaters induce film-evaporation at the fluid interface to generate thrust. FEMTA has been in development at Purdue University since 2015 under the NASA SmallSat Technology Partnership Program and is currently on its 5th generation design. While these generations of FEMTA have successfully demonstrated the viability of the propulsion technique under ideal conditions, multiple reliability and performance related issues have been identified. More specifically, high vacuum tests have shown that the current FEMTA design is susceptible to quiescent propellant mass loss due to ice generation and leaking at the nozzle. These mass ejections can limit the lifespan and performance of the thruster and can induce undesired attitude perturbations on the host spacecraft. The purpose of this researchidentify the root causes of the quiescent mass loss mechanims hrough simulation and direct experimentation. Based on the results of these investigations, a next generation design is proposed, fabricated, and tested. Microfabrication was performed at Purdue’s Birck Nanotechnology Center and vacuum and thrust stand tests were performed at the High Vacuum Lab in the Aerospace Sciences Laboratory at Purdue.</p></div></div></div>

<div>High speed flow problems of practical interest require a solution of nonequilibrium aerothermochemistry to accurately predict important flow phenomena including surface heat transfer and stresses. As a majority of these flow problems are in the continuum regime, Computational Fluid Dynamics (CFD) is a useful tool for flow modeling. This work presents the development of a nonequilibrium add-on solver to ANSYS Fluent utilizing user-defined-functions to model salient aspects of nonequilibrium flow in air. The developed solver was verified for several benchmark nonequilibrium flow problems and compared with the available experimental data and other nonequilibrium flow simulations. <br></div><div><br></div><div>The rate of dissociation behind a strong shock in thermochemical nonequilibrium depends on the vibrational excitation of molecules. The Macheret-Fridman (MF) classical impulsive model provides analytical expressions for nonequilibrium dissociation rates. The original form of the model was limited to the dissociation of homonuclear molecules. In this work, a general form of the MF model has been derived and present macroscopic rates applicable for modeling dissociation in CFD. Additionally, some improvements to the prediction of mean energy removed in dissociation in the MF-CFD model has been proposed based on the comparisons with available QCT data. In general, the results from the MF-CFD model upon investigating numerous nonequilibrium flows are promising and the model shows a possibility of becoming the standard tool for investigating nonequilibrium flows in CFD.</div><div><br></div><div>The aerodynamic deorbit experiment (ADE) CubeSat has dragsail to accompany accelerated deorbiting of a CubeSat post-mission. A good estimation of the aerothermal load on a reentry CubeSat is paramount to ensure a predictable reentry. This study investigates the aerothermal load on an ADE CubeSat using the direct simulation Monte Carlo (DSMC) methods and Navier-Stokes-Fourier continuum based methods with slip boundary conditions. The aerothermal load on an ADE CubeSat at 90 km altitude from the DSMC and continuum methods were consistent with each other. The continuum breakdown at a higher altitude of 95 km resulted in a strong disagreement between the continuum and DSMC solutions. Overall, the continuum methods could offer a considerable computational cost saving to the DSMC methods in predicting aerothermal load on an ADE CubeSat at low altitudes.<br> </div>

This master thesis focuses on literature research of problematics linked to power actuators working on electromagnetic principle to accelerate metal projectiles. It’s goal is mathematical analysis and constuction of selected type of electromagntic power actuator – induction coilgun

This master thesis focuses on literature research of problematics linked to power actuators working on electromagnetic principle to accelerate metal projectiles. It’s goal is mathematical analysis and constuction of selected type of electromagntic power actuator – induction coilgun

This master thesis deals with the design of a handbike with electric propulsion. The final design is created in regard to knowledge from design and technical analysis and also to desficiencies of contemporary products. The purpose of this work is to create design of electrohandbike which will respect user and his needs from ergonomical and esthetic point of view.

A reusable oxidizer-cooled annular aerospike nozzle was designed for testing on a labscale PMMA-N20[1] hybrid rocket motor at Cal Poly-SLO.[2] The detailed design was based on the results of previous research involving cold-flow testing of annular aerospike nozzles and hot-flow testing of oxidizer-cooled converging-diverging nozzles. In the design, nitrous oxide is routed to the aerospike through a tube that runs up the middle of the combustion chamber. The solid fuel is arranged in an annular configuration, with a solid cylinder of fuel in the center of the combustion chamber and a hollow cylinder of fuel lining the circumference of the combustion chamber. The center fuel grain insulates the coolant from the heat of the combustion chamber. The two-phase mixture of nitrous oxide then is routed through channels that cool the copper surface of the aerospike. The outer copper shell is brazed to a stainless steel core that provides structural rigidity. The gaseous N2O flows from the end of aerospike to provide base bleed, compensating for the necessary truncation of the spike. Sequential and fully-coupled thermal-mechanical finite element models developed in Abaqus CAE were used to analyze the design of the cooled aerospike. The stress and temperature distributions in the aerospike were predicted for a 10-sec burn time of the hybrid rocket motor. [1] PMMA stands for polymethyl methacrylate, a thermoplastic commonly known by the brand name Plexiglas®. N2O is the molecular formula for nitrous oxide. [2] California Polytechnic State University, San Luis Obispo

The circulation control system onboard Cal Poly's Advanced Model for Extreme Lift and Improved Aeroacoustics was a critical component of a highly complex wind tunnel model produced in order to fulfill the requirements of a NASA Research Announcement awarded to David Marshall of the Aerospace Engineering Department. The model was based on a next generation, 150 passenger, regional, cruise efficient, short take-off and landing concept aircraft that achieved high lift through circulation control wings and over-the-wing mounted engines. The wind tunnel model was 10-ft in span, used turbine propulsion simulators, and had a functioning circulation control system driven from tunnel supplied high pressure air. Wind tunnel test results will be compiled into an open-source database intended for validation of predictive tools whose purpose is to advance the state- of-the-art in predictive capabilities for the next generation aircraft configurations. The model's circulation control system produced highly directional, nonuniform flow, and required significant modification in order to generate flow suitable for representation in predictive software. The effort and methods used to generate uniform flow along the circulation control slots is detailed herein. Additionally the results of the system characterization are presented and include a thorough analysis of the slot height, the wing symmetry, and total pressure at the circulation control jet exit. These datasets are intended to aid in making adjustments to the simulation such that it accurately reflects the condition at which the model was tested. Many flow visualization results from the wind tunnel test are also presented to serve as a medium of comparison for results from predictive tools. Oil flow visualization was conducted at many test conditions and provides insight to AMELIA's surface flow in blown and unblown regions. Of particular interest were streamlines at the wingblend, which exhibited some outboard turning, and streamlines on the lower surface where the leading edge stagnation point was investigated. Smoke flow visualization was also utilized to explore the flowfield. The deflection of a individual streamline, under the influence of a changing discharge coefficient as investigated along with the discharge coefficients effect on the extended flowfield. Collectively, the images depict the massive augmentation of the flowfield caused by the presence of the circulation control wing.

The subject of this diploma thesis is the design solution of the axial headstock, its gear box and main drive. This headstock is used for heavy duty gantry type machines from the production of TOS Kurim company. The thesis contains technical solution of the headstock, an analysis of the construction of the headstock drive, overview of the possible propulsion units, choice of the best drive variant, control calculations, technical proposal of drive design and economical evaluation of the whole reconstruction.

This thesis deals with the construction of an active transtibial prosthesis for recreational swimming. Its aim is to create a working sample of such a prosthesis. The device will primarily serve the patient to move in water - it will be attached to the stump and it will be driven forward. Secondarily, the prosthesis can be used for walking, in the sense that the patient attaches the prosthesis near to a water surface (pool, lake, ocean) and comes with it to the shore. As part of the design and electronics it will be necessary to determine the characteristics of the human resistance in water depending on speed and to determine the necessary propulsion (thrust) strength of the propeller, which will serve as the driving force. Subsequently there will be formed structural variants and selected the best one. This will become default for design and construction of simplified testing device, which will be tested for static thrust. After verifying the functionality and reliability, the final functional sample will be manufactured and tested too.

The future of aviation technology is transitioning to cleaner, more efficient and higher endurance aircraft solutions. As fully electric propulsion systems still fall short of the operational requirements of modern day aircraft, there is increasing pressure and demand for the aviation industry to explore alternatives to fossil fuel driven propulsion systems. The primary focus of this research is to experimentally evaluate hybrid electric propulsion systems (HEPS) for Unmanned Aerial Vehicles (UAV) which combine multiple power sources to improve performance. HEPS offer several potential benefits over more conventional propulsion systems such as a smaller environmental impact, lower fuel consumption, higher endurance and novel configurations through distributed propulsion. Advanced operating modes are also possible with HEPS, increasing the vehicle’s versatility and redundancy in case of power source failure. The primary objective of the research is to combine all of the components of a small-scale HEPS together in a modular test bench for evaluation. The test bench uses components sized for a small-scale UAV including a 2.34kW two-stroke 35cc engine and a 1.65kW brushless DC motor together with an ESC capable of regenerative braking. Individual components were first tested to characterize performance, and then all components were assembled together in a parallel configuration to observe system-level performance. The parallel HEPS is capable of functioning in the four required operating modes: EM Only, ICE Only, Dash Mode (combined EM and ICE power) as well as Regenerative Mode where the onboard batteries get recharged. Further, the test bench was implemented with a supervisory controller to optimize system performance and run each component in the most efficient region to achieve torque requirements programmed into mission profiles. The logic based controller operates with the ideal operating line (IOL) concept and is implemented with a custom LabView GUI. The system is able to run on electric power or ICE power interchangeably without making any modifications to the transmission as the one-way bearing assembly engages for whichever power source is rotating at the highest speed. The most impressive of these sets of tests is the Dash mode testing where the output torque of the propeller is supplied from both the EM and ICE. Working in tandem, it was proved that the EM was drawing 19.9A of current which corresponds to an estimated 0.57Nm additional torque to the propeller for a degree of hybridization of 49.91%. Finally, the regenerative braking mode was proven to be operational, capable of recharging the battery systems at 13A. All of these operating modes attest to the flexibility and convenience of having a hybrid-electric propulsion system. The results collected from the test bench were validated against the models created in the aircraft simulation framework. This framework was created in MATLAB to simulate the performance of a small UAV and compare the performance by swapping in various propulsion systems. The purpose of the framework is to make direct comparisons of HEPS performance for parallel and series architectures against conventional electric and gasoline configuration UAVs, and explore the trade-offs. Each aircraft variable in the framework was modelled parametrically so that parameter sweeps could be run to observe the impact on the aircraft’s performance. Finally, rather than comparing propulsion systems in steady-state, complex mission profiles were created that simulate real life applications for UAVs. With these experiments, it was possible to observe which propulsion configurations were best suited for each mission type, and provide engineers with information about the trade-offs or advantages of integrating hybrid-electric propulsion into UAV design. In the Pipeline Inspection mission, the exact payload capacities of each aircraft configuration could be observed in the fuel burn versus CL,cruise parameter sweep exercise. It was observed that the parallel HEPS configuration has an average of 3.52kg lower payload capacity for the 35kg aircraft (17.6%), but has a fuel consumption reduction of up to 26.1% compared to the gasoline aircraft configuration. In the LIDAR Data collection mission, the electric configuration could be suitable for collection ranges below 100km but suffers low LIDAR collection times. However, at 100km LIDAR collection range, the series HEPS has an endurance of 16hr and the parallel configuration has an endurance of 19hr. In the Interceptor mission, at 32kg TOW, the parallel HEPS configuration has an endurance/TOW of 1.3[hr/kg] compared to 1.15[hr/kg] for the gasoline aircraft. This result yields a 13% increase in endurance from 36.8hr for gasoline to 41.6hr for the parallel HEPS. Finally, in the Communications Relay mission, the gasoline configuration is recommended for all TOW above 28kg as it has the highest loiter endurance. / Graduate