Dissertations / Theses: 'Aerospace Engineering - Propulsion'

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Zhu, Dawei. "Supercirculation Aerodynamic-Propulsion Test Rig Instrumentation Development." Ohio University / OhioLINK, 2006. http://rave.ohiolink.edu/etdc/view?acc_num=ohiou1142542776.

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2

Gilpin, Matthew R. "High temperature latent heat thermal energy storage to augment solar thermal propulsion for microsatellites." Thesis, University of Southern California, 2016. http://pqdtopen.proquest.com/#viewpdf?dispub=10160163.

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Solar thermal propulsion (STP) offers an unique combination of thrust and efficiency, providing greater total ΔV capability than chemical propulsion systems without the order of magnitude increase in total mission duration associated with electric propulsion. Despite an over 50 year development history, no STP spacecraft has flown to-date as both perceived and actual complexity have overshadowed the potential performance benefit in relation to conventional technologies. The trend in solar thermal research over the past two decades has been towards simplification and miniaturization to overcome this complexity barrier in an effort finally mount an in-flight test.

A review of micro-propulsion technologies recently conducted by the Air Force Research Laboratory (AFRL) has identified solar thermal propulsion as a promising configuration for microsatellite missions requiring a substantial Δ V and recommended further study. A STP system provides performance which cannot be matched by conventional propulsion technologies in the context of the proposed microsatellite ''inspector" requiring rapid delivery of greater than 1500 m/s ΔV. With this mission profile as the target, the development of an effective STP architecture goes beyond incremental improvements and enables a new class of microsatellite missions.

Here, it is proposed that a bi-modal solar thermal propulsion system on a microsatellite platform can provide a greater than 50% increase in Δ V vs. chemical systems while maintaining delivery times measured in days. The realization of a microsatellite scale bi-modal STP system requires the integration of multiple new technologies, and with the exception of high performance thermal energy storage, the long history of STP development has provided "ready" solutions.

For the target bi-modal STP microsatellite, sensible heat thermal energy storage is insufficient and the development of high temperature latent heat thermal energy storage is an enabling technology for the platform. The use of silicon and boron as high temperature latent heat thermal energy storage materials has been in the background of solar thermal research for decades without a substantial investigation. This is despite a broad agreement in the literature about the performance benefits obtainable from a latent heat mechanisms which provides a high energy storage density and quasi-isothermal heat release at high temperature.

In this work, an experimental approach was taken to uncover the practical concerns associated specifically with applying silicon as an energy storage material. A new solar furnace was built and characterized enabling the creation of molten silicon in the laboratory. These tests have demonstrated the basic feasibility of a molten silicon based thermal energy storage system and have highlighted asymmetric heat transfer as well as silicon expansion damage to be the primary engineering concerns for the technology. For cylindrical geometries, it has been shown that reduced fill factors can prevent damage to graphite walled silicon containers at the expense of decreased energy storage density.

Concurrent with experimental testing, a cooling model was written using the "enthalpy method" to calculate the phase change process and predict test section performance. Despite a simplistic phase change model, and experimentally demonstrated complexities of the freezing process, results coincided with experimental data. It is thus possible to capture essential system behaviors of a latent heat thermal energy storage system even with low fidelity freezing kinetics modeling allowing the use of standard tools to obtain reasonable results.

Finally, a technological road map is provided listing extant technological concerns and potential solutions. Improvements in container design and an increased understanding of convective coupling efficiency will ultimately enable both high temperature latent heat thermal energy storage and a new class of high performance bi-modal solar thermal spacecraft.

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Eilers, Shannon Dean. "Development of the Multiple Use Plug Hybrid for Nanosats (Muphyn) Miniature Thruster." DigitalCommons@USU, 2013. https://digitalcommons.usu.edu/etd/1726.

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The Multiple Use Plug Hybrid for Nanosats (MUPHyN) prototype thruster incorporates solutions to several major challenges that have traditionally limited the deployment of chemical propulsion systems on small spacecraft. The MUPHyN thruster offers several features that are uniquely suited for small satellite applications. These features include 1) a non-explosive ignition system, 2) non-mechanical thrust vectoring using secondary fluid injection on an aerospike nozzle cooled with the oxidizer flow, 3) a non-toxic, chemically-stable combination of liquid and inert solid propellants, 4) a compact form factor enabled by the direct digital manufacture of the inert solid fuel grain. Hybrid rocket motors provide significant safety and reliability advantages over both solid composite and liquid propulsion systems; however, hybrid motors have found only limited use on operational vehicles due to 1) difficulty in modeling the fuel flow rate 2) poor volumetric efficiency and/or form factor 3) significantly lower fuel flow rates than solid rocket motors 4) difficulty in obtaining high combustion efficiencies. The features of the MUPHyN thruster are designed to offset and/or overcome these shortcomings. The MUPHyN motor design represents a convergence of technologies, including hybrid rocket regression rate modeling, aerospike secondary injection thrust vectoring, multiphase injector modeling, non-pyrotechnic ignition, and nitrous oxide regenerative cooling that address the traditional challenges that limit the use of hybrid rocket motors and aerospike nozzles. This synthesis of technologies is unique to the MUPHyN thruster design and no comparable work has been published in the open literature.

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Collie, Wallis Vernon. "Design and Analysis of an Unmanned Aerial Vehicle Propulsion System with Fluidic Flow Control Inside a Highly Compact Serpentine Inlet Duct." NCSU, 2003. http://www.lib.ncsu.edu/theses/available/etd-11282003-145453/.

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The benefits of highly compact serpentine inlet ducts extend from reductions in overall aircraft weight to higher survivability, as well as allow the aircraft designer greater flexibility in propulsion system integration. Unfortunately, due to the extreme wall curvature, these ducts result in significant flow distortion and total pressure losses at the engine face. It has been shown that active flow control in the form of micro-fluidic vortex generators significantly helps to reduce these losses. To date, these systems have only been tested in a laboratory setting in which items such as flow control air supply, system and subsystem size, weight, and location are not major factors. Subscale unmanned aerial vehicles provide a real world test bed to help overcome these constraints at a lower cost and lower risk as compared to full scale aircraft testing. This work presents the design, integration, testing, and analysis of an unmanned aerial vehicle?s propulsion system that implements fluidic flow control inside a highly compact serpentine inlet duct in order to reduce engine face distortion and increase propulsion system performance.

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Armstrong, Isaac W. "Development and Testing of Additively Manufactured Aerospike Nozzles for Small Satellite Propulsion." DigitalCommons@USU, 2019. https://digitalcommons.usu.edu/etd/7428.

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Automatic altitude compensation has been a holy grail of rocket propulsion for decades. Current state-of-the-art bell nozzles see large performance decreases at low altitudes, limiting rocket designs, shrinking payloads, and overall increasing costs. Aerospike nozzles are an old idea from the 1960’s that provide superior altitude-compensating performance and enhanced performance in vacuum, but have survivability issues that have stopped their application in satellite propulsion systems. A growing need for CubeSat propulsion systems provides the impetus to study aerospike nozzles in this application. This study built two aerospike nozzles using modern 3D metal printing techniques to test aerospikes at a size small enough to be potentially used on a CubeSat. Results indicated promising in-space performance, but further testing to determine thermal limits is deemed necessary.

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Marklund, Hanna. "Supersonic Retro Propulsion Flight Vehicle Engineering of a Human Mission to Mars." Thesis, Luleå tekniska universitet, Rymdteknik, 2019. http://urn.kb.se/resolve?urn=urn:nbn:se:ltu:diva-75820.

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A manned Mars mission will require a substantial increase in landed mass compared to previous robotic missions, beyond the capabilities of current Entry Descent and Landing, EDL, technologies, such as blunt-body aeroshells and supersonic disk-gap-band parachutes. The heaviest payload successfully landed on Mars to date is the Mars Science Laboratory which delivered the Curiosity rover with an approximate mass of 900 kg. For a human mission, a payload of magnitude 30-50 times heavier will need to reach the surface in a secure manner. According to the Global Exploration Roadmap, GER, a Human Mission to Mars, HMM, is planned to take place after year 2030. To prepare for such an event several technologies need maturing and development, one of them is to be able to use and accurately asses the performance of Supersonic Retro Propulsion, SRP, another is to be able to use inflatable heat shields. This internal study conducted at the European Space Agency, ESA, is a first investigation focusing on the Entry Descent and Landing, EDL, sequence of a manned Mars lander utilising an inflatable heatshield and SRP, which are both potential technologies for enabling future landings of heavy payloads on the planet. The thesis covers the areas of aerodynamics and propulsion coupled together to achieve a design, which considers the flight envelope constraints imposed on human missions. The descent has five different phases and they are defined as circular orbit, hypersonic entry, supersonic retropropulsion, vertical turn manoeuvre and soft landing. The focus of this thesis is on one of the phases, the SRP phase. The study is carried out with the retro-thrust profile and SRP phase initiation Mach number as parameters. Aerodynamic data in the hyper and supersonic regime are generated using Computational Fluid Dynamics, CFD, to accurately assess the retropropulsive performance. The basic concept and initial sizing of the manned Mars lander builds on a preliminary technical report from ESA, the Mission Scenarios and Vehicle Design Document. The overall optimisation process has three parts and is based on iterations between the vehicle design, CFD computations in the software DLR-Tau and trajectory planning in the software ASTOS. Two of those parts are studied, the vehicle design and the CFD,to optimise and evaluate the feasibility of SRP during the descent and test the design parameters of the vehicle. This approach is novel, the efficiency and accuracy of the method itself is discussed and evaluated. Initially the exterior vehicle Computer Aided Design, CAD, model is created, based on the Mission Scenarios and Vehicle Design Document, however updated and furthered. The propulsion system is modelled and evaluated using EcosimPRO where the nozzle characteristics, pressure levels and chemistry are defined, and later incorporated in the CAD model. The first iteration of the CFD part has an SRP range between Mach 7 and 2, which results in an evaluation of five points on the trajectory. The thrust levels, the corresponding velocity, altitude and atmospheric properties at those points can then be evaluated and later incorporated in ASTOS. ASTOS, in turn, can simulate the full trajectory from orbit to landing including the CFD data of the SRP phase. Due to time limitation only one iteration of the vehicle design and the SRP range was completed. However, the goals of the study were reached. A first assessment of SRP in Mars atmosphere has been carried out, and the aerodynamic and propulsive data has been collected to be built on in the future. The results indicate that the engines can start at a velocity of Mach 7. They also show consistency with similar studies conducted in Earths atmosphere. The current vehicle design, propulsion system and SRP range can now be furthered, updated and advanced in order to optimise the different descent phases in combination with future results from ASTOS.

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Connolly, Joseph. "Aero-Propulso-Elastic Analysis of a Supersonic Transport." The Ohio State University, 2018. http://rave.ohiolink.edu/etdc/view?acc_num=osu1543337967878799.

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Chamberlain, Britany L. "Additively-Manufactured Hybrid Rocket Consumable Structure for CubeSat Propulsion." DigitalCommons@USU, 2018. https://digitalcommons.usu.edu/etd/7285.

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Three-dimensional, additive printing has emerged as an exciting new technology for the design and manufacture of small spacecraft systems. Using 3-D printed thermoplastic materials, hybrid rocket fuel grains can be printed with nearly any cross-sectional shape, and embedded cavities are easily achieved. Applying this technology to print fuel materials directly into a CubeSat frame results in an efficient, cost-effective alternative to existing CubeSat propulsion systems. Different 3-D printed materials and geometries were evaluated for their performance as propellants and as structural elements. Prototype "thrust columns" with embedded fuel ports were printed from a combination of acrylonitrile utadiene styrene (ABS) and VeroClear, a photopolymer substitute for acrylic. Gaseous oxygen was used as the oxidizer for hot-fire testing of prototype thrusters in ambient and vacuum conditions. Hot-fire testing in ambient and vacuum conditions on nine test articles with a combined total of 25 s burn time demonstrated performance repeatability. Vacuum specific impulse was measured at over 167 s and maximum thrust of individual thrust columns at 9.5 N. The expected ΔV to be provided by the four thrust columns of the consumable structure is approximately 37 m/s. With further development and testing, it is expected that the consumable structure has the potential to provide a much-needed propulsive solution within the CubeSat community with further applications for other small satellites.

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9

Bertuzzi, Alberto. "Microcontroller based flow control for spacecraft electric propulsion." Master's thesis, Alma Mater Studiorum - Università di Bologna, 2018.

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10

Cheney, Liam Jon. "Development of Safety Standards for CubeSat Propulsion Systems." DigitalCommons@CalPoly, 2014. https://digitalcommons.calpoly.edu/theses/1180.

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The CubeSat community has begun to develop and implement propulsion systems. This movement represents a new capability which may satisfy mission needs such as orbital and constellation maintenance, formation flight, de-orbit, and even interplanetary travel. With the freedom and capability granted by propulsion systems, CubeSat providers must accept new responsibilities in proportion to the potential hazards that propulsion systems may present. The Cal Poly CubeSat program publishes and maintains the CubeSat Design Specification (CDS). They wish to help the CubeSat community to safety and responsibly expand its capabilities to include propulsive designs. For this reason, the author embarked on the task of developing a draft of safety standards CubeSat propulsion systems. Wherever possible, the standards are based on existing documents. The author provides an overview of certain concepts in systems safety with respect to the classification of hazards, determination of required fault tolerances, and the use of inhibits to satisfy fault tolerance requirements. The author discusses hazards that could exist during ground operations and through launch with respect to hazardous materials and pressure systems. Most of the standards related to Range Safety are drawn from AFSPCMAN 91-710. Having reviewed a range of hypothetical propulsion system architectures with an engineer from Range Safety at Vandenberg Air Force Base, the author compiled a case study. The author discusses many aspects of orbital safety. The author discusses the risk of collision with the host vehicle and with third party satellites along with the trackability of CubeSats using propulsion systems. Some recommendations are given for working with the Joint Functional Component Command for Space (JFCC SPACE), thanks to the input of two engineers who work with the Joint Space Operations Center (JSpOC). Command Security is discussed as an important aspect of a mission which implements a propulsion system. The author also discusses End-of-Life procedures such as safing and de-orbit operations. The orbital safety standards are intended to promote “good citizenship.” The author steps through each proposed standard and offers justification. The author is confident that these standards will set the stage for a dialogue in the CubeSat community which will lead to the formulation of a reasonable and comprehensive set of standards. The author hopes that the discussions given throughout this document will help CubeSat developers to visualize the path to flight readiness so that they can get started on the right foot.

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Schoeffler, Lara Elaine. "Orbital Dynamics of Space Nuclear Propulsion Systems." Case Western Reserve University School of Graduate Studies / OhioLINK, 2021. http://rave.ohiolink.edu/etdc/view?acc_num=case1618332162764726.

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Ramunno, Michael Angelo. "Control Optimization of Turboshaft Engines for a Turbo-electric Distributed Propulsion Aircraft." The Ohio State University, 2020. http://rave.ohiolink.edu/etdc/view?acc_num=osu1587657623577243.

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Rojas, Sigala Mauro. "Study of Launcher Recovery Systems." Thesis, Luleå tekniska universitet, Rymdteknik, 2020. http://urn.kb.se/resolve?urn=urn:nbn:se:ltu:diva-80722.

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The space sector has been evolving due to the fast-technological advancements generating a reduction of manufacturing, cost and size in space missions, where highly capable performing small satellites are becoming the standard in this industry. Furthermore, the high launching cost limits the trend of cost reduction for the space missions, since the small satellites are sent as a second payload. An alternative to reduce this limitation is using reusable launchers which are key in the future of space industry, once they are optimized in efficiency and reliability. Therefore, an opportunity of design is presented, since the increase of small satellites missions requires a reduction of the cost in launch services a suitable option for the future market are the reusable launchers. The problematic of using recovery systems and reuse parts of the vehicle is the increase of weight due to the added systems that the vehicle needs to be recovered. This paper presents different engines and calculate the performance of each engine based on the needs of missions for small satellites. The starting conditions will be that the payload needs to be launched in low circular or elliptical orbits (altitudes of between 300 and 650 km) and the engine has the ability of vertical take-off, vertical landing. The design will also take into account the possibility of reusing parts of the vehicle and the reentry capability. Different combination of engines and fuels are setup in various configurations. For each case the mass analysis will be developed which will allow to calculate the performance for each engine. The important parameters are the number and type of engines, the ratios of the masses, the thrust-to-weight ratio and specific impulse. Once the mass analysis is obtained the following procedure is the selection of the design considering the empty mass. The best combination of characteristics of the engines will be the suitable candidate. Different assistance systems and techniques for the recovery are assessed to obtain a suitable option to improve the efficiency. The expected results are the calculation of the engine performance and how the selected design can be suitable for the space launcher sector for the small satellites. The expected results are a feasible vehicle for small satellites design based in the calculation of the engine parameters together with an efficient launch recovery system. The conclusion is that the space sector can benefit from the design, demonstrating that a launch vehicle with the reusable characteristics

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Abada, Hashim H. "Turboelectric Distributed Propulsion System for NASA Next Generation Aircraft." Wright State University / OhioLINK, 2017. http://rave.ohiolink.edu/etdc/view?acc_num=wright1515501052742277.

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Chakravarthula, Venkata Adithya. "Transient Analysis of a Solid Oxide Fuel Cell/ Gas Turbine Hybrid System for Distributed Electric Propulsion." Wright State University / OhioLINK, 2016. http://rave.ohiolink.edu/etdc/view?acc_num=wright1484651177170392.

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Brezina, Aron Jon. "Measurement of Static and Dynamic Performance Characteristics of Electric Propulsion Systems." Wright State University / OhioLINK, 2012. http://rave.ohiolink.edu/etdc/view?acc_num=wright1340066274.

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Urban, Peter J. "Non-Intrusive Optical Measurement of Electron Temperature in Near Field Plume of Hall Thruster." Cleveland State University / OhioLINK, 2018. http://rave.ohiolink.edu/etdc/view?acc_num=csu1528216243044845.

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Pansolin, Denis. "An Experimental and Numerical Study of High Temperature Gaseous Flow through an Open Cell Silicon Carbide Foam Heater." ScholarWorks@UNO, 2019. https://scholarworks.uno.edu/td/2702.

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Grannan, Nicholas D. "Design and Structural Analysis of a Dual Compression Rotor." University of Dayton / OhioLINK, 2013. http://rave.ohiolink.edu/etdc/view?acc_num=dayton1366644139.

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Bacciaglia, Antonio. "Design and Development of a Propulsion System for a Water-Air Unmanned Vehicle." Master's thesis, Alma Mater Studiorum - Università di Bologna, 2018.

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This work aims to contribute on the design of a Bimodal Unmanned Underwater and Air System (BUUAS). This research project will first present background informations of current hybrid UAV concepts with a focus on the different types of propulsion mechanisms used in air/water transition. Then a brief description of BUUAS will lead to requirements for the transition mechanism that this work aims to develop. After a section dedicated to the description of a short-impulse thruster layout, theoretical and experimental approaches are used to determine the amount of thrust generated. As second design step, a simplified UAV version is used to test the transition phase using the designed thruster. Finally, a section is dedicated to a design layout description with the thruster and an optimized propeller. Future work is proposed to continue in the development of this project, with a short description of folding wing and propulsion system integration concepts.

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Sergent, Aaronn. "Optimal Sizing and Control of Battery Energy Storage Systems for Hybrid-Electric, Distributed-Propulsion Regional Aircraft." The Ohio State University, 2020. http://rave.ohiolink.edu/etdc/view?acc_num=osu1595519141013663.

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Belapurkar, Rohit K. "Stability and Performance of Propulsion Control Systems with Distributed Control Architectures and Failures." The Ohio State University, 2013. http://rave.ohiolink.edu/etdc/view?acc_num=osu1357309068.

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Hartwig, Jason W. "Liquid Acquisition Devices for Advanced In-Space Cryogenic Propulsion Systems." Case Western Reserve University School of Graduate Studies / OhioLINK, 2014. http://rave.ohiolink.edu/etdc/view?acc_num=case1396562473.

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Wilson, Matthew D. "Catalytic Decomposition of Nitrous Monopropellant for Hybrid Motor Ignition." DigitalCommons@USU, 2013. https://digitalcommons.usu.edu/etd/1496.

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Nitrous oxide (N2O) is an inexpensive and readily available non-toxic rocket motor oxidizer. It is the most commonly used oxidizer for hybrid bipropellant rocket systems, and several bipropellant liquid rocket designs have also used nitrous oxide. In liquid form, N2O is highly stable, but in vapor form it has the potential to decompose exothermically, releasing up to 1865 Joules per gram of vapor as it dissociates into nitrogen and oxygen. Consequently, it has long been considered as a potential "green" replacement for existing highly toxic and dangerous monopropellants. This project investigates the feasibility of using the nitrous oxide decomposition reaction as a monopropellant energy source for igniting liquid bipropellant and hybrid rockets that already use nitrous oxide as the primary oxidizer. Because nitrous oxide is such a stable propellant, the energy barrier to dissociation is quite high; normal thermal decomposition of the vapor phase does not occur until temperatures are above 800 C. The use of a ruthenium catalyst decreases the activation energy for this reaction to allow rapid decomposition below 400 C. This research investigates the design for a prototype device that channels the energy of dissociation to ignite a laboratory scale hybrid rocket motor.

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Deans, Matthew Charles. "The Simulation, Development, and Testing of a Staged Catalytic Microtube Ignition System." Case Western Reserve University School of Graduate Studies / OhioLINK, 2013. http://rave.ohiolink.edu/etdc/view?acc_num=case1343318716.

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Gonzalez, Marin Victor Alberto. "Methods to operate and evaluate the performance of a cold-gas CubeSat propulsion system on a magnetically stabilised satellite." Thesis, KTH, Rymdteknik, 2020. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-288714.

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Propulsion systems allow satellites to perform many functionalities in space, such as orbital station keeping, reentry control, attitude control, orbital transferring, rendezvous operation, and even more thrilling, interplanetary travel. Indeed, propulsion systems in satellites have fostered a new favourable era of space exploration and application, therefore, detailed processes to operate propulsion systems need to be developed so that space missions, carrying this valuable system, are completed successfully. The aim of this study is to describe the most relevant operating procedures for the cold gas propulsion system NanoProp 3U, developed by GomSpace, on-board the 3U CubeSat MIST satellite developed by KTH. Procedures, such as power levels, telemetry considerations, propellant mass determination, Fault Detection Isolation and Recovery analysis, and decommissioning plan allow proper operation of NanoProp according to the mission requirements determined for MIST mission. Moreover, this study describes detailed mission experiments to be performed with NanoProp with the objective of assessing the performance delivered by the propulsion system itself, and other on-board subsystems which are required for monitoring and controlling the spacecraft according to the effects generated by the propulsion system. The planning and operation of a propulsion system should be outlined on-ground, during the mission design, so a clear understanding of the characteristics and limitations of the system are highlighted towards the development of a secure and solid space mission.
Framdrivningssystem tillåter satelliter att utföra många funktioner i rymden, som t.ex. att hålla konstant avstånd till en annan rymdfarkost, utlösa återinträde i atmosfären, attitydstyrning, manövrera mellan olika omloppsbanor, och, till och med, interplanetära uppdrag. Framdrivningssystem i satelliter har främjat en ny lovande era av rymdforskning och praktisk tillämpning av rymden, och därför behöver detaljerade, men praktiskt hanterbara, metoder för att operativt använda framdrivningssystem utvecklas. Basen för detta arbete är att beskriva de mest relevanta driftsrutinerna för framdrivningssystemet NanoProp 3U, utvecklat av GomSpace, för användning ombord på MIST-satelliten (en 3U Cubesat) som utvecklats av KTH. Aspekter på NanoProps användning i MIST som förbrukning av elektrisk energi, telemetribehov, drivmedelsmassa, hantering av felfunktioner (upptäckt och avhjälpande) och avveckling av satelliten vid drifttidens slut analyseras i detalj. Dessutom analyserar detta arbete hur detaljerade driftprov kan utföras med NanoProp i syfte att bedöma de prestanda som framdrivningssystemet tillhandahåller och hur dessa prov påverkar och stöds av driften av satellitens övriga delsystem. Det övergripande syftet med detta arbete är således att utveckla en metod för att planera driften av ett framdrivningssystem under ett satellitprojekts definitions- och utvecklingsfaser så att en tydlig förståelse av systemets egenskaper och begränsningar leder till ett säkert och stabilt rymduppdrag.

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D'Anniballe, Alessandro. "Development of a sizing tool for preliminary mission analysis and design of propulsion systems for orbit control of small satellites in LEO -VLEO." Master's thesis, Alma Mater Studiorum - Università di Bologna, 2017. http://amslaurea.unibo.it/14719/.

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The ever-growing necessity for faster and cheaper access to space, makes the building of artificial satellites to shift more and more towards small measures. The most exemplar case of small satellites is given by the so called CubeSats, i.e. modular satellites practically ‘built in blocks’ of approximately the same size and weight. Such an approach allows to fasten the design and decrease the overall project complexity, but at the same time has many limitations. The main one is equipping the satellites with a propulsion system for the control of their operative orbit. Such a task normally requires a consistent amount of propellant, and so a specific dedicated propulsion system, that weighs consistently (in terms of both mass and volume) on the budget of these small spacecrafts. The present thesis studies the feasibility of providing small satellites with a propulsion system that would enable them to perform orbit control maneuvers all along the mission duration. The concept is to create a computer tool able to carry out a rapid analysis of the satellite mission, for the determination of the needed Δv, and then a preliminary design of the main components of the required propulsion system. Different propulsion technologies can in this way be considered, being then able to do a trade-off to choose the best solution, in terms of mass and performances. Satellite models ranging from nano to mini-sat standard in LEO-VLEO missions of different durations (2, 5 and 7 years) have been used for feasibility simulations, and re-sults show that the use of some propulsion technology is possible to reach the fixed mission goals.

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Plank, Jack R. "Nuclear Thermal Propulsion Cool-Down Phase Optimization Through Quasi-Steady Computational Analysis, and the Effect of Auxiliary Heat Removal Systems." The Ohio State University, 2021. http://rave.ohiolink.edu/etdc/view?acc_num=osu1618934609976051.

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Scharlemann, Carsten A. "Investigation of thrust mechanisms in a water fed pulsed plasma thruster." The Ohio State University, 2003. http://rave.ohiolink.edu/etdc/view?acc_num=osu1070354149.

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Romano, Federico. "Q1D unsteady ballistic model for solid rocket motors performance prediction." Master's thesis, Alma Mater Studiorum - Università di Bologna, 2021.

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The simulation tool ROBOOST, in use at the Alma Propulsion Lab of the University of Bologna – Forlì Campus, exploits a hybrid ballistic model 0D-1D. The need of a complete Q1D model for the entire combustion time, from motor start-up to burn out arised. The present work is devoted to the development and test of a Q1D unsteady ballistic model for solid rocket motors performance prediction. The newly developed code, called SOL1D, is written in Matlab environment and is capable of predicting the time and space evolution of all the main thermodynamic variables during the solid rocket motor combustion process. The model has been tested and validated on a BARIA motor, thus demonstrating its adherence to experimental data. SOL1D paves the way for future works aimed at simulating performances of actual launchers.

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Bodnar, Maxwell J. "The Creation, Analysis, and Verification of a Comprehensive Model of a Micro Ion Thruster." DigitalCommons@CalPoly, 2015. https://digitalcommons.calpoly.edu/theses/1565.

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A computational model of the micro-ion thruster MiXI has been developed, analyzed, and partially verified. This model includes submodels that govern the physical, magnetic, electrostatic, plasma physics, and power deposition of the thruster. Over the past few years, theses have been conducted with the goal of running tests and analyzing the results; this model is used to understand how the thruster components interact so as to make predictions about, and allow for optimization of, the thruster operation. Testing is then performed on the thruster and the results are compared to the output of the code. The magnetic structure of the thruster was analyzed and numerous different configurations generated which were also evaluated by the optimizer and tested. Using the different configurations, models, and optimization tools, the total efficiency of the thruster is theoretically able to reach 69.4%. Operational testing of the thruster at many different throttle settings demonstrated a maximum total efficiency of 45.9 ±24.6%, discharge loss values as low as 109 ±25 eV/ion, and total power required as low as 50.5 ±0.1W to maintain thruster operation with beam extraction. Measurements of the plasma were taken using a Langmuir probe and the interpretation of the tests are used to verify the plasma physics submodel. Power draw measurements and analysis of the throttle inputs during testing are compared to the performance model outputs but were not accurate or consistent enough to fully verify the power deposition and plasma physics models. Analysis of the models and operational testing in this study have led to an increased understanding of the performance and operation of the MiXI-CP-V3 thruster, furthering the effort to create an efficient, flight capable micro-ion thruster.

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Bulut, Jane. "Design and CFD analysis of the demonstrator aerospike engine for a small satellite launcher application." Master's thesis, Alma Mater Studiorum - Università di Bologna, 2020.

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Starting with a brief overview of thrust generation for launchers, this study focuses on the design process of the demonstrator aerospike engine, DEMOP-1, of the Pangea Aerospace's commercial grade engine and its flow field analysis. The primary goal of the study is to obtain the plug nozzle design delivers 30 kN thrust using cryogenic liquid oxygen (LOX) as the oxidizer and cryogenic liquid methane (LCH4) as the fuel, with the mixture ratio of 3.4. Design parameters considered as 30 bar of combustion chamber pressure (Po) and expansion ratio as 15 for an optimum expanded nozzle. On the basis of decided design characteristics, Angelino's method is used to design the nozzle contour through MATLAB. The flow field over the aerospike analyzed using commercial CFD program FLUENT for sea level, optimum expansion and vacuum conditions. Flow simulations are carried out for air (specific heat ratio, gamma= 1.4), and afterwards based on the obtained thrust values at each altitude for air, expected thrust values for the real propellant, LOX/LCH4 (specific heat ratio, gamma = 1.1664), are calculated. Finally, the study is concluded with the comparison of trend in thrust and specific impulse for conventional bell nozzle and aerospike. For the conventional bell engine the values obtained in commercial computational simulation of chemical rocket propulsion and combustion software RPA for bell nozzle with same characteristics with aerospike, Po = 30 bar and expansion ratio = 15, are taken as reference for sea level, optimum expansion level and vacuum condition performance. Due to its ability to adopt the altitude, aerospike delivers higher performance at the low altitudes with respect to the conventional bell nozzle which has the same expansion ratio and combustion chamber pressure. Last in order but not in importance, after obtaining the flow field on plug of the aerospike, the shock wave impingement on the nozzle surface at sea level has been investigated.

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Ball, Tyler M. "CFD as Applied to the Design Of Short Takeoff and Landing Vehicles Using Circulation Control." DigitalCommons@CalPoly, 2008. https://digitalcommons.calpoly.edu/theses/50.

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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.

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McCrink, Matthew H. "Development of Flight-Test Performance Estimation Techniques for Small Unmanned Aerial Systems." The Ohio State University, 2015. http://rave.ohiolink.edu/etdc/view?acc_num=osu1449142886.

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Staniscia, Giada. "Development of a Low Earth Orbit Mission Preliminary Analysis Tool." Thesis, KTH, Rymdteknik, 2019. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-265613.

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The objective of this project is the development of a mission analysis tool for the nanosatellite company GomSpace Sweden. Although there are many existing software, they can be quite complicated and time consuming to use. The goal of this work is to build a simple app to be used at the earliest stages of space missions in order to obtain key figures of merit quickly and easily. By comparing results, assessing the feasibility of customer needs, analysing how various parameters affect each other, it enables immediate deeper understanding of the implications of the main design decisions that are taken at the very beginning of a mission. The tool shall aid the system engineering process of determining orbit manoeuvre capability specifically for CubeSat electric propulsion systems taking into account the most relevant factors for perturbation in Low Earth Orbit (LEO), i.e. atmospheric drag and Earth’s oblateness effects. The manoeuvres investigated are: orbit raising from an insert orbit to an operating orbit, orbit maintenance, deorbiting within the space debris mitigation guidelines and collision avoidance within the 12 to 24 hours that the system has to react. The manoeuvres cost is assessed in terms of Delta v requirements, propellant mass and transfer times. The tool was developed with MATLAB and packaged as a standalone Linux application.
Målet med detta examensarbete var att utveckla ett verktyg för missionsanalys för nanosatellitföretaget GomSpace Sweden. Det finns många andra mjukvaror för att nå samma mål men de är ofta komplicerade och tidskrävande. Det specifika målet var således att skapa en enkel applikation som kan användas i de tidiga stegen av utformning av rymduppdrag för att snabbt och enkelt få fram viktiga parametrar. Genom att jämföra resultat, uppskatta genomförbarheten av kundbehov och analysera hur olika parametrar påverkar varandra kan omedelbar förståelse erhållas rörande påverkan av designbeslut som tas i början av rymduppdragen. Verktyget ska stödja systemingenjörsprocessen genom att uppskatta banförflyttningskapacitet för elektriska framdrivningssystem för CubeSats och ta i beaktande de mest relevanta faktorerna gällande störningar i låg jordbana (LEO), i.e. atmosfäriskt motstånd och effekterna av Jordens form. De undersökta manövrarna är: banhöjning från injektionsbana till operationell bana, banunderhåll, bansänkning som följer riktlinjerna för rymdskrot och kollisionsundvikande inom de 12 till 24 timmar som systemet har på sig att reagera. Kostnaden för manövrarna är uppskattade genom DeltaV-krav, massan av bränslet och förflyttningstider. Verktyget utvecklades med MATLAB och paketerades som en fristående applikation i Linux.

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Claypool, Ian Randolph. "A theoretical and numerical study of the use of grid embedded axial magnetic fields to reduce charge exchange ion induced grid erosion in electrostatic ion thrusters." Columbus, Ohio : Ohio State University, 2007. http://rave.ohiolink.edu/etdc/view?acc%5Fnum=osu1172690635.

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Capatina, Allen A. C. "AXISYMMETRIC BI-PROPELLANT AIR AUGMENTED ROCKET TESTING WITH ANNULAR CAVITY MIXING ENHANCEMENT." DigitalCommons@CalPoly, 2015. https://digitalcommons.calpoly.edu/theses/1493.

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Performance characterization was undertaken for an air augmented rocket mixing duct with annular cavity configurations intended to produce thrust augmentation. Three mixing duct geometries and a fully annular cavity at the exit of the nozzle were tested to enable thrust comparisons. The rocket engine used liquid ethanol and gaseous oxygen, and was instrumented with sensors to output total thrust, mixing duct thrust, combustion chamber pressure, and propellant differential pressures across Venturi flow measurement tubes. The rocket engine was tested to thrust maximum, with three different mixing ducts, three major combustion pressure sets, and a nozzle exit plane annular cavity (a grooved ring). The combustion pressures tested were , , and allowing for a nozzle pressure ratio range of relative to ambient pressure. The mixture ratio was fuel rich throughout all tests. The engine operated very consistently throughout all the tests performed; however, pressure losses in the feed system prevented higher combustion pressures from being tested. Three mixing ducts of the same outer diameter were tested. The short and diverging ducts were the same length and the long duct was long. The short and long ducts created positive mixing duct thrust and the diverging duct created negative mixing duct thrust. The long duct case did show better performance than the no duct case when the total thrust was divided by combustion pressure and nozzle throat area. The long duct always created several times more mixing duct thrust than either the short or diverging ducts, but none of the mixing ducts created positive overall thrust augmentation in the over expanded cases tested. The mixing duct thrusts ranged between and . As the combustion pressures were increased, getting closer the nozzle’s optimal expansion, the mixing duct thrusts started converging indicating a difference between nozzle operation at over expanded and under expanded. The annular cavity had a noticeable effect on the thrust of the engine and the appearance of the plume. The total thrust of the system was decreased by a maximum of and the plume was more sharply defined when the annular cavity was attached. Better mixing between the primary (engine exhaust) flow and the secondary (ambient air) flow was promoted by the annular cavity because it increased the shear layer’s turbulence and the increased turbulence reduced thrust. The greater mixing also allowed for secondary combustion which made the plumes more sharply defined. The annular cavity was also seen to enhance the mixing duct thrusts for all three mixing ducts.

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Persson, Robert. "PPS5000 Thruster Emulator Architecture Development & Hardware Design." Thesis, Luleå tekniska universitet, Rymdteknik, 2019. http://urn.kb.se/resolve?urn=urn:nbn:se:ltu:diva-72827.

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This Master's Thesis handles prestudy work and early hardware development that resulted in architectural definitions and prototype hardware of electronic ground support equipment. This equipment is destined to emulate the electric power consumption of the PPS5000 Hall Effect Thruster (HET), for use in satellite end-to-end tests of the all-electric Geostationary Satellite Electra, developed at OHB Sweden AB. The Thruster Emulator (TEM) was defined through a resulting compilation of intricate interdependent components that interface the satellite power system and the thruster, which yielded an architecture development to support some basic predefined emulator requirements. This architecture was then analyzed to form a base-line conceptual function of the emulator system, which incorporates the entire HET functionality. Six primary HET impedances were defined, of which the three most complex impedances were investigated fully. For the primary thruster discharge, research is shown of the complexity of implementing advanced electronic load hardware directly to the satellite's 5kW power system with respect to the transient primary plasma discharge during thruster start up, and with limitations on the electronic load reducing emulator-thruster similarities. Additionally, a fully functional plasma ignition emulator prototype circuit board was built to be used in the final hardware of the TEM to emulate the external HET cathode start-up functionality. Finally, a feasibility study for designing a possible solution for the large PPS5000 electromagnet impedance was performed, resulting in the manufacture of two prototype inductors with unsatisfying performance results according to the design requirements.

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Pereira, Roger Michael. "The I2T5 : Enhancement of the Thermal Design of an Iodine Cold Gas Thruster." Thesis, Luleå tekniska universitet, Rymdteknik, 2020. http://urn.kb.se/resolve?urn=urn:nbn:se:ltu:diva-80702.

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The I2T5, an iodine-propelled, cold gas thruster, developed by ThrustMe, France, is the first of its kind to make it successfully to space. Due to its simple, reliable and cost-effective design, it is a suitable propulsion system for CubeSat missions with low delta-V (ΔV) requirements. To ensure that the I2T5 performs at its peak, it is crucial to maintain good thermal control of the thruster, to keep it within the operational temperature range. The first flight measurements of the I2T5 provided insight into its thermal performance. It was observed that the required temperature to sublimate the iodine propellant was not reached within the expected time frame, which led to a longer warm-up period, and a reduction in thrust. The problem arose due to an unforeseen conductive thermal contact between the tank and the thruster walls. This thesis delves deeper into this issue, and focuses on alleviating the total conductive heat loss from the tank to the satellite frame, where the I2T5 is integrated. The insulating washer-bolt configuration of the I2T5 side panels is observed to be responsible for the conductive heat transfer. A preliminary analysis is performed to obtain an initial maximum for the conductive heat flux lost to the satellite frame. A plan of action is then determined to optimise the geometry, material or configuration of the insulating washers to lower the maximum heat flux value. Following this, an experiment was conducted with a new washer-bolt configuration to determine the heat flux values. A case study is performed for the orbital environment heat fluxes that the I2T5 would receive if it were integrated to a CubeSat in sun-synchronous orbit. An overview of results shows that, for the thermal simulations, all the methods employed to reduce the conductive heat loss at the frame were effective. The experiment provided neutral results, and would need to be repeated with different experimental parameters to have a clear perspective of the heat losses. In reality, the satellite frame receives radiative fluxes in addition to conductive heat fluxes, but radiation is not considered for this thesis, and is suggested as a prospective study.

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Girardello, Carlo. "Optical Analysis of Plasma : Flame Emission in Cryogenic Rocket Engines." Thesis, Luleå tekniska universitet, Rymdteknik, 2019. http://urn.kb.se/resolve?urn=urn:nbn:se:ltu:diva-76097.

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This thesis contains the results of optical flame emission measurements of the Vulcain 2.1engine and the plasma emission spectroscopy of the Lumen Project engine. The plume spectroscopyis analyzed, ordered and studied in detail to offer the best possible molecular composition.The main focus relied on the hydroxide radical, blue radiation and other moleculesanalysis of the intensities encountered during the tests. The plasma emission spectroscopy isfocused on the determination of the plasma temperature value in LIBS measurements. Thehydrogen plasma temperature determination of the local thermodynamic equilibrium, followedby the carbon and sequentially oxygen plasma is obtained. The quality of the LTE isto be determined to judge the truthworthness of the determined temperatures. Both the testsare analyzed thanks to the use of spectrographs, cameras and dedicated software for opticalapplications. The results related to the Vulcain 2.1 LOX/LH2 engine showed the evolutionof the plume in different ROF or pressure variations. Furthermore, the results of the LumenProject LOX/methane engine led to the determination of the plasma temperatures and a firstestimation of the LTE quality.
Die vorliegende Arbeit präsentiert die Ergebnisse der Abgasstrahlspektroskopie des H2/LOXVulcain 2.1 Triebwerks und der Zündplasma Spektroskopie des CH4/LOX Triebwerks desLUMEN Projektes. Die Abgasstrahlspektroskopie wurde analysiert und im Detail untersuchtum die am besten passende molekulare Zusammensetzung herauszuarbeiten. DasHauptaugenmerk liegt dabei auf dem Hydroxyl- Radikal, der Blauen Strahlung und molekularerIntensitätsanalyse. Bei der Zündplasmaanalyse liegt der Fokus auf der Bestimmungdes LTE Zustands (Lokales thermodynamisches Gleichgewicht) in LIBS. Die Temperaturdes Wasserstoff-, Kohlenstoff und Sauerstoffplasmas wird herangezogen, um die Qualitätdes LTE Zustands zu beurteilen. Für die Testdurchführung wurden Spektrographen, Kamerasund bestimmte Auswertungstools für optische Anwendungen benutzt. Das Verhaltendes Vulcain 2.1 Abgasstrahls abhängig von verschiedenen ROF und Druckstufen ist in denErgebnissen beschrieben. Für das LUMEN Triebwerk konnten erste Zündplasmatemperaturenbestimmt werden und geben einen Rückschluss auf die Qualität des LTE.

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Gilles, Paul M. "Performance Enhancement and Characterization of an Electromagnetic Railgun." DigitalCommons@CalPoly, 2019. https://digitalcommons.calpoly.edu/theses/2107.

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Collision with orbital debris poses a serious threat to spacecraft and astronauts. Hypervelocity impacts resulting from collisions mean that objects with a mass less than 1g can cause mission-ending damage to spacecraft. A means of shielding spacecraft against collisions is necessary. A means of testing candidate shielding methods for their efficacy in mitigating hypervelocity impacts is therefore also necessary. Cal Poly’s Electromagnetic Railgun was designed with the goal of creating a laboratory system capable of simulating hypervelocity (≥ 3 km/s) impacts. Due to several factors, the system was not previously capable of high-velocity (≥ 1 km/s) tests. A deficient projectile design is revised, and a new design is tested. The new projectile design is demonstrated to enable far greater performance than the previous design, with a muzzle velocity ≥ 1 km/sbeing verified during testing, and an energy conversion efficiency of 2.7%. A method of improving contact and controlling wear at the projectile/rail interface using silver plating and conductive silver paste is validated. A mechanism explaining the problem of internal arcing within the railgun barrel is proposed, and design recommendations are made to eliminate arcing on the basis of the work done during testing. The primary structural members are found to be deficient for their application and a failure analysis of a failed member, loading analysis of the railgun barrel, and design of new structures is undertaken and presented.

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Yentsch, Robert J. "Three-Dimensional Shock-Boundary Layer Interactions in Simulations of HIFiRE-1 and HIFiRE-2." The Ohio State University, 2013. http://rave.ohiolink.edu/etdc/view?acc_num=osu1384195671.

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Morrisey, Bryan J. "Multidisciplinary Design Optimization of an Extreme Aspect Ratio HALE UAV." DigitalCommons@CalPoly, 2009. https://digitalcommons.calpoly.edu/theses/113.

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ABSTRACT Multidisciplinary Design Optimization of an Extreme Aspect Ratio HALE UAV Bryan J. Morrisey Development of High Altitude Long Endurance (HALE) aircraft systems is part of a vision for a low cost communications/surveillance capability. Applications of a multi payload aircraft operating for extended periods at stratospheric altitudes span military and civil genres and support battlefield operations, communications, atmospheric or agricultural monitoring, surveillance, and other disciplines that may currently require satellite-based infrastructure. Presently, several development efforts are underway in this field, including a project sponsored by DARPA that aims at producing an aircraft that can sustain flight for multiple years and act as a pseudo-satellite. Design of this type of air vehicle represents a substantial challenge because of the vast number of engineering disciplines required for analysis, and its residence at the frontier of energy technology. The central goal of this research was the development of a multidisciplinary tool for analysis, design, and optimization of HALE UAVs, facilitating the study of a novel configuration concept. Applying design ideas stemming from a unique WWII-era project, a “pinned wing” HALE aircraft would employ self-supporting wing segments assembled into one overall flying wing. The research effort began with the creation of a multidisciplinary analysis environment comprised of analysis modules, each providing information about a specific discipline. As the modules were created, attempts were made to validate and calibrate the processes against known data, culminating in a validation study of the fully integrated MDA environment. Using the NASA / AeroVironment Helios aircraft as a basis for comparison, the included MDA environment sized a vehicle to within 5% of the actual maximum gross weight for generalized Helios payload and mission data. When wrapped in an optimization routine, the same integrated design environment shows potential for a 17.3% reduction in weight when wing thickness to chord ratio, aspect ratio, wing loading, and power to weight ratio are included as optimizer-controlled design variables. Investigation of applying the sustained day/night mission requirement and improved technology factors to the design shows that there are potential benefits associated with a segmented or pinned wing. As expected, wing structural weight is reduced, but benefits diminish as higher numbers of wing segments are considered. For an aircraft consisting of six wing segments, a maximum of 14.2% reduction in gross weight over an advanced technology optimal baseline is predicted.

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Benyo, Theresa L. "Analytical and Computational Investigations of a Magnetohydrodynamic (MHD) Energy-Bypass System for Supersonic Turbojet Engines to Enable Hypersonic Flight." Kent State University / OhioLINK, 2013. http://rave.ohiolink.edu/etdc/view?acc_num=kent1369153719.

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Choi, Jinbae. "Closed-Loop Optimal Control of Discrete-Time Multiple Model Linear Systems with Unknown Parameters." Case Western Reserve University School of Graduate Studies / OhioLINK, 2016. http://rave.ohiolink.edu/etdc/view?acc_num=case1441178373.

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Kummer, Joseph. "Simulation of the cross-flow fan and application to a propulsive airfoil concept." Related electronic resource: Current Research at SU : database of SU dissertations, recent titles available full text, 2006. http://proquest.umi.com/login?COPT=REJTPTU0NWQmSU5UPTAmVkVSPTI=&clientId=3739.

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Rhodes, Gregory D. "Experimental Investigations of the Propulsive Fuselage Concept." The Ohio State University, 2018. http://rave.ohiolink.edu/etdc/view?acc_num=osu1523355985000317.

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Plewacki, Nicholas. "Modeling High Temperature Deposition in Gas Turbines." The Ohio State University, 2020. http://rave.ohiolink.edu/etdc/view?acc_num=osu1587714424017527.

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Yezeguelian, Axel. "Modelling and Simulation of a Propulsive Hybridisation for a Light Fixed-wing Aircraft." Thesis, KTH, Flygdynamik, 2019. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-261222.

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Propulsive hybridisation fits in with the sustainable development policies of many companies which are part of the transportation industry. Actually, it makes it possible either to reduce fuel consumption or to improve the aircraft performance at a fixed fuel burn. However, the current technologies of batteries restrain a more regular use in light aviation. For this project this issue is confirmed as both the quasi-static performance assessment and the dynamic studies show that the endurance objective cannot be improved with Li/Ion batteries. However, it is possible to act directly on the engine performance by placing a thermal energy recovery system on exhaust gas pipes to take advantage of their high temperatures, greatly boosting the aircraft performance in cruise.
Hybridisering av framdrivningssystem passar in med hållbar utvecklingspolitik av många företag inom transportbranschen. Faktiskt tillåter det antingen att minska bränsleförbrukningen eller att förbättra flygplans prestandorna. Ändå är aktuella batteritekniken fortfarande ett problem för en mer frekvent användning av hybridisering förlättflyget. För det här projektet bekräftas det eftersom kvasistatiska och dynamiska studierna visar att hybridsystemet med aktuella Li/Ion batterierna förbättrar inte flygdurationen. Ändå är det möjligt att ingripa på motorprestandorna genom att återanvända värmeenergin av avgas för att förbättra flygdurationen.

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Dlima, Kendrick M. "Conceptual Design of a South Pole Carrier Pigeon UAV." DigitalCommons@CalPoly, 2020. https://digitalcommons.calpoly.edu/theses/2145.

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Currently, the South Pole has a large data problem. It is estimated that 1.2 TB of data is being produced every day, but less than 500 GB of that data is being uploaded via aging satellites to researchers in other parts of the world. This requires those at the South Pole to analyze the data and carefully select the parts to send, possibly missing out on vital scientific information. The South Pole Carrier Pigeon will look to bridge this data gap. The Carrier Pigeon will be a small unmanned aerial vehicle that will carry a 30 TB solid-state hard drive from the South Pole to various destinations in the Southern Hemisphere, but it has been designed to y to Christchurch, New Zealand. This 87 lb. UAV will be able to y 3,650 nmi. up to 25,000 ft., using a 5.7 hp. engine. It will feature an de-icing system on the leading edge of its 8 ft. span wing to allow it to y through cold, moist climates. It will have a 39 in. long fuselage with a tail boom of 33 in. The aircraft has been designed to be made out of composites, thus reducing both the weight of the aircraft as well as its drag. It has been designed to come apart in order to be shipped successfully to the South Pole. There, it will be assembled and launched via a custom pneumatic launcher. It will y autonomously to 15,000 ft. and cruise climb throughout the flight to 25,000 ft., before descending to its destination. There, it will be caught by a net restraint system, where the hard drive will be extracted. The Carrier Pigeon is truly a unique vehicle for its size, range, and robustness.

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Dissertations / Theses on the topic 'Aerospace Engineering - Propulsion'

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