Dissertations / Theses: 'Aerospike engines'

1

Korak, Ghosh. "Model predictive control for civil aerospace gas turbine engines." Thesis, University of Sheffield, 2013. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.595827.

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Kiker, Adam Paul. "Experimental Investigations of Mini-Pulsejet Engines." NCSU, 2005. http://www.lib.ncsu.edu/theses/available/etd-08112005-134914/.

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An experimental 8 cm pulsejet was developed using scaling laws from research on both 50 and 15 cm pulsejets. The 8 cm jet operates in three different inlet configurations?conventional, perpendicular, and rearward. The rearward configuration features inlets facing in the opposite direction of the flight path and develops the maximum net thrust. Using a high frequency pressure transducer, the operational frequency of the pulsejet was obtained by monitoring the combustion chamber pressure. It was found that in the rearward configuration, the operational frequency of the jet decreases with increasing inlet length. In addition, the combustion chamber peak pressure rise per cycle increases significantly if the exhaust diameter is reduced. Using information from the 8 cm pulsejet, a 4.5 cm pulsejet was developed and is operational.

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Schoen, Michael Alexander. "Experimental Investigations in 15 Centimeter Class Pulsejet Engines." NCSU, 2005. http://www.lib.ncsu.edu/theses/available/etd-08082005-095911/.

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Testing is performed on the 15 centimeter class pulsejet engine in order to develop, study, and explore the operational characteristics. Valved and valveless operation, hydrogen and propane fuels, various fuel injection methods, and a range of geometric configurations are investigated for operational feasibility. The scaling capabilities of a valveless 15 centimeter class pulsejet of conventional design are studied by methodically varying inlet length, exit length, exit geometry, and inlet area to combustor area ratio (Ai/Ac). Engine performance is defined by measuring chamber pressure, internal gas temperatures, time-resolved thrust, operational frequency, and fuel flow rate. The scaling capability is characterized by the success of self-sustained combustion for each corresponding geometric configuration. Tail pipe length is found to be a function of valveless inlet length and may be further minimized by the addition of a diverging exit nozzle. Chemical kinetic times and Ai/Ac prove to be the two prominent controlling parameters in determining scaling behavior.

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OVERMAN, NICHOLAS. "FLAMELESS COMBUSTION APPLICATION FOR GAS TURBINE ENGINES IN THE AEROSPACE INDUSTRY." University of Cincinnati / OhioLINK, 2006. http://rave.ohiolink.edu/etdc/view?acc_num=ucin1163776616.

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Siddiqi, Majid. "Turbine Engine Control and Diagnostics." The Ohio State University, 2005. http://rave.ohiolink.edu/etdc/view?acc_num=osu1420210543.

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Reichel, Jonathan R. "Parametric study of liquid fuel jet in crossflow at conditions typical of aerospace applications." Thesis, Atlanta, Ga. : Georgia Institute of Technology, 2008. http://hdl.handle.net/1853/22590.

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Nilamdeen, Mohamed Shezad. "An uncoupled multiphase approach towards modeling ice crystals in jet engines." Thesis, McGill University, 2010. http://digitool.Library.McGill.CA:8881/R/?func=dbin-jump-full&object_id=92185.

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Pietroniro, Asuka Gabriele. "Modelling coaxial jets relevant to turbofan jet engines." Thesis, KTH, Mekanik, 2016. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-200909.

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Simulations of subsonic turbulent coaxial hot jets were conducted on two types ofunstructured grids within the framework of STAR-CCM+. The study case is based on atypical airliner turbofan engine model with a core nozzle and a fan nozzle, having a bypassratio of five. The two meshes used are a polyhedral one, suitable for complex surfaces, and atrimmed one mainly made of hexahedral cells. The sensitivity of the study case to variousinputs is attested using second and third order upwind schemes, modelling turbulence with aSST k-omega model. The project proves to be a valid feasibility study for a steady-statesolution on which an aeroacoustic analysis could be based in future works.

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Cosher, Christopher R. "Detailed Analysis of Previous Data Relevant to Foreign Particle Ingestion by GasTurbine Engines and Application to Modern Engines." The Ohio State University, 2016. http://rave.ohiolink.edu/etdc/view?acc_num=osu1461152408.

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10

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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Staniszewski, Marcin. "Simulation of tri-axially braided composites half-cylinder behavior during balistic [sic] impact." Akron, OH : University of Akron, 2007. http://rave.ohiolink.edu/etdc/view?acc%5Fnum=akron1177978645.

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Thesis (M.S.)--University of Akron, Dept. of Civil Engineering, 2007.
"May, 2007." Title from electronic thesis title page (viewed 4/28/2009) Advisor, Wieslaw K. Binienda; Committee members, Craig C. Menzemer, Ala Abbas; Department Chair, Wieslaw K. Binienda; Dean of the College, George K. Haritos; Dean of the Graduate School, George R. Newkome. Includes bibliographical references.

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Hoopes, Daniel (Daniel Michael). "Improving producibility in aerospace engine manufacturing : process automation vs. process reengineering." Thesis, Massachusetts Institute of Technology, 2008. http://hdl.handle.net/1721.1/43836.

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Thesis (M.B.A.)--Massachusetts Institute of Technology, Sloan School of Management; and, (S.M.)--Massachusetts Institute of Technology, Dept. of Materials Science and Engineering; in conjunction with the Leaders for Manufacturing Program at MIT, 2008.
Includes bibliographical references (p. 43).
In any aerospace manufacturing operation, including Pratt & Whitney's Compression Systems Module Center, producibility problems can be major drivers of cost. Much of the literature focuses on design for manufacturability as a solution to producibility problems. While this is a valuable approach, this study focuses on manufacturing process improvement as a solution to producibility issues. Two methods of process improvement are discussed, process automation and process reengineering. This thesis first surveys some of the major producibility problems at Pratt & Whitney's Compression Systems Module Center, as well as some of the efforts underway to address them. One of the largest issues, operator data input errors, is described in detail as a case study. Wireless gauging with automatic offset adjustment is proposed as a focused technological solution to this issue. As part of this study, funding has been obtained to implement the solution and testing has been conducted. This is an example of process automation. However, a broader process reengineering effort is also proposed. The fundamental question of why producibility problems tend to persist is also examined.
by Daniel Hoopes.
S.M.
M.B.A.

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13

Whitfield, Clifford A. "Experimental Development and Investigation of Propeller / Jet Engine Interactions." The Ohio State University, 2006. http://rave.ohiolink.edu/etdc/view?acc_num=osu1421160400.

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14

Roberts, James W. "Further calculations of the performance of turbofan engines incorporating a wave rotor." Thesis, Monterey, California : Naval Postgraduate School, 1990. http://handle.dtic.mil/100.2/ADA240867.

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Thesis (M.S. in Aeronautical Engineering)--Naval Postgraduate School, September 1990.
Thesis Advisor(s): Shreeve, Raymond P. Second Reader: Hobson, Garth V. "September 1990." Description based on title screen as viewed on December 18, 2009. DTIC Descriptor(s): Rotors, Turbofan Engines, Waves, Gases, Pressure, Ratios, Computer Programs, Cycles. DTIC Identifier(s): Wave Rotors, Rotors, Waves, Theses. Author(s) subject terms: Turbofan Engines, Turbofan engines with a Wave Rotor. Includes bibliographical references (p. 95-96). Also available in print.

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15

Andrén, Hugo. "Towards Zero Defects in the Aerospace Industry through Statistical Process Control : A Case Study at GKN Aerospace Engine Systems." Thesis, Luleå tekniska universitet, Institutionen för ekonomi, teknik och samhälle, 2020. http://urn.kb.se/resolve?urn=urn:nbn:se:ltu:diva-79927.

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With the ongoing transformation of modern manufacturing systems in an industry 4.0 environment, industrial actors may see great improvements with respect to quality towards a state of near zero defects. For the aerospace industry, where increased quality and reduced risk is strongly related, new technologies may be used in manufacturing to see to the increasing demands on products. The safety, as well as the manufacturing complexity of products and processes, make the collected measurement data an integral asset for enterprises within the aerospace industry. Collected data may be analysed using statistical tools and methods to improve process capability and, in extension, product quality. Communicating the need for zero defects, original equipment manufacturers demand increased capability from product and component manufacturers. Hence, zero defects are typically operationalised through exhibiting a process capability of Cpk= 2.0. In response to the challenge, GKN Aerospace need to raise the traditional process capability targets of Cpk=1.33. By employing an exploratory research strategy with a deductive approach, the thesis combines theoretical knowledge from the literature with empirical findings in a thematic analysis. The thematic analysis was conducted by employing six phases as suggested by Braun and Clarke (2006) and resulted in the identification of 50 codes from a total of 459 data extracts. Based on the empirical interview data, a framework for how zero defects is interpreted at GKN Aerospace was developed, which describes zero defects as a cycle. Taking into account that zero defects is operationalised through Cpk= 2.0, the cycle consists of six phases that start with a vision and is completed by delivering a true and reliable Cpk of 2.0. In addition, the codes from the thematic analysis were collated into a thematic mind map, focusing on key aspects of working with statistical process control (SPC) to support zero defects. Two main themes are presented in the mind map, statistical approach to improvement work; highlighting necessary aspects of statistical process control and measurability, and removing barriers for improvement; highlighting fundamental organisational barriers that impede proactive quality improvement. To support the findings and give a practical example of how process data may be presented and analysed using tools and methods within statistical process control, an SPC study was conducted on a set of data. In the SPC study, the construction and analysis of individuals Shewhart control charts and moving range charts were described in detail. These procedures provide better insights about process behaviour through statistical thinking and thus better knowledge on how to approach more proactive process improvements.

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Ha, Dong Keun. "Reducing drag of a commuter train, using engine exhaust momentum." Thesis, California State University, Long Beach, 2016. http://pqdtopen.proquest.com/#viewpdf?dispub=10108178.

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The objective of this thesis was to perform numerical investigations of two different methods of injecting fluid momentum into the air flow above a commuter train to reduce its drag. Based on previous aerodynamic modifications of heavy duty trucks in improving fuel efficiency, two structural modifications were designed and applied to a Metrolink Services commuter train in the Los Angeles (LA) County area to reduce its drag and subsequently improve fuel efficiency. The first modification was an L-shaped channel, added to the exhaust cooling fan above the locomotive roof to divert and align the exhaust gases in the axial direction. The second modification was adding an airfoil shaped lid over the L-shape channel, to minimize the drag of the perturbed structure, and thus reduce the overall drag.

The computational fluid dynamic (CFD) software CCM+ from CD-Adapco with the ?-? turbulence model was used for the simulations. A single train set which consists of three vehicles: one locomotive, one trailer car and one cab car were used. All the vehicles were modeled based on the standard Metrolink fleet train size. The wind speed was at 90 miles per hour (mph), which is the maximum speed for the Orange County Metrolink line. Air was used as the exhaust gas in the simulation. The temperature of the exhausting air emitting out of the cooling fan on the roof was 150 F and the average fan speed was 120 mph.

Results showed that with the addition of the lid, momentum injection results in reduced flow separation and pressure recovery behind the locomotive, which reduces the overall drag by at least 30%.

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Zinnecker, Alicia M. "Modeling for Control Design of an Axisymmetric Scramjet Engine Isolator." The Ohio State University, 2012. http://rave.ohiolink.edu/etdc/view?acc_num=osu1354215841.

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18

Léonard, Pauline. "Sustainability assessment of composites in aero-engine components." Thesis, Luleå tekniska universitet, Institutionen för teknikvetenskap och matematik, 2019. http://urn.kb.se/resolve?urn=urn:nbn:se:ltu:diva-75369.

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Environmental issues such as climate change are leading to important sustainability challenges in the aerospace industry. Composites are light materials that are extensively used to replace metals and reduce the aircraft weight, the goal being to decrease the fuel consumption in flight and limit the emission of greenhouse gases. However, these high performance materials are associated with a complex supply chain including energy-consuming processes. Most of the decommissioned composite products are currently landfilled and nothing proves that the weight reduction allowed by these materials compensates those negative aspects. The purpose of this master thesis is to determine if the introduction of composites in aero-engines can be sustainable and how it can be achieved. To do so, three polymer-matrix composite components from GKN Aerospace have been studied and compared with their metallic baseline from environmental, social and economic perspectives. Several options for materials selection, manufacturing processes and recycling possibilities have been investigated in the same way. The assessment on GKN Aerospace’s components showed that the weight savings provided by composites have a strong and positive influence on their sustainability. Component B shows the best results: with 16% of weight savings with composites versus the titanium baseline, it appears clearly that the composite version is the most sustainable one. Component A2 composite version also provides interesting weight savings (14%) but has an aluminum baseline, which makes the composite component more sustainable in some aspects but not all of them, especially economically speaking. Finally, for component A1, the composite version, which does not provide weight savings, is more economically feasible, but quite tight with the titanium baseline on environmental and social aspects. Therefore, it appears that composite components are more likely to be sustainable if they provide significant weight reduction and if the baseline is titanium. A few strategies would merit attention to make future composite components more sustainable. On the one hand, using thermoplastic composites have potential to reduce the environmental, social and economic impact. In fact, these materials can be fully recycled and reused, present less risks to handle and can be produced for a lower cost. Nevertheless, the knowledge on these materials is more limited than on thermoset composite and the implementation of such a solution will take time. On the second hand, introducing composite recycling processes in the products lifecycle can increase a lot the sustainability of composite components. The manufacturing scrap and the decommissioned products can both be recycled in order to reduce the environmental impact and generate benefits by re-using or selling the recycled material.

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Miquel, Valentin. "Propellant Feeding System of a Liquid Rocket With Multiple Engines." Thesis, KTH, Rymdteknik, 2020. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-276460.

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Multiengine first stages are the new trend in recent rockets. Reusability and an oxygen/methane based engine complete this picture. ArianeGroup wants to develop its own rocket following these principles. This thesis presents the study of the feeding system for a seven Prometheus engine rocket. Several ways of connecting propellant tanks to engines were proposed and analyzed. Two configurations were selected and studied with more detail. One consists of a main feeding line which is then split in seven secondary lines. The other one adds one rank of pipes to reduce the number of feeding valves. Their performances were assessed according to classic space industry drivers. Furthermore, the impact of the two solutions on the efficiency of the tank was evaluated. CAD drawings and simulation models were made and could be a base for future work if one of the systems is chosen. The study shows that a falcon 9 like feeding system is performant in terms of mass and pressure losses but another cost-effective configuration is possible and gives good results.
Första stegen med flera motorer är den nya trenden i de senaste raketerna. Återanvändbart och en syre och metan-baserad motor kompletterar denna bild. ArianeGroup vill utveckla sin egen raket enligt dessa principer. Denna avhandling presenterar studien av drivmedelsrör för en sju Prometheus-motorraket. Flera sätt att ansluta drivmedelstankar till motorer föreslogs och analyserades. Två konfigurationer valdes ut och studerades mer detaljerat. En består av en huvudlinje som sedan delas upp i sju sekundära linjer som på SpaceX Falcon 9. Den andra lösningen lägger till en rang av rör för att minska antalet ventiler. Deras prestanda utvärderades först enligt klassiska kriterier för rymdindustrin. Dessutom utvärderades de två lösningarnas påverkan på tankens effektivitet. CAD-ritningar och simuleringsmodeller gjordes och kan vara en bas för framtida arbeten om ett av systemen väljs. Studien visar att ett Falcon 9-liknande konfiguration har bättre prestanda när det gäller massa och tryckförluster men en annan kostnadseffektiv konfiguration är möjlig och ger goda resultat.

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Buettner, Robert W. "Dynamic Modeling and Simulation of a Variable Cycle Turbofan Engine with Controls." Wright State University / OhioLINK, 2017. http://rave.ohiolink.edu/etdc/view?acc_num=wright1496179248257409.

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Salamon, Nicholas C. "Analysis of Nuclear Thermal Rocket Engine Coolant Channel Designs Enabled byAdditive Manufacturing." The Ohio State University, 2020. http://rave.ohiolink.edu/etdc/view?acc_num=osu1587590263667569.

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22

Walter, S. F. "Optimization of pressure probe placement and data analysis of engine-inlet distortion." Thesis, University of Colorado at Boulder, 2017. http://pqdtopen.proquest.com/#viewpdf?dispub=10244536.

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The purpose of this research is to examine methods by which quantification of inlet flow distortion may be improved upon. Specifically, this research investigates how data interpolation effects results, optimizing sampling locations of the flow, and determining the sensitivity related to how many sample locations there are. The main parameters that are indicative of a "good" design are total pressure recovery, mass flow capture, and distortion. This work focuses on the total pressure distortion, which describes the amount of non-uniformity that exists in the flow as it enters the engine. All engines must tolerate some level of distortion, however too much distortion can cause the engine to stall or the inlet to unstart. Flow distortion is measured at the interface between the inlet and the engine.

To determine inlet flow distortion, a combination of computational and experimental pressure data is generated and then collapsed into an index that indicates the amount of distortion. Computational simulations generate continuous contour maps, but experimental data is discrete. Researchers require continuous contour maps to evaluate the overall distortion pattern. There is no guidance on how to best manipulate discrete points into a continuous pattern. Using one experimental, 320 probe data set and one, 320 point computational data set with three test runs each, this work compares the pressure results obtained using all 320 points of data from the original sets, both quantitatively and qualitatively, with results derived from selecting 40 grid point subsets and interpolating to 320 grid points. Each of the two, 40 point sets were interpolated to 320 grid points using four different interpolation methods in an attempt to establish the best method for interpolating small sets of data into an accurate, continuous contour map. Interpolation methods investigated are bilinear, spline, and Kriging in Cartesian space, as well as angular in polar space. Spline interpolation methods should be used as they result in the most accurate, precise, and visually correct predictions when compared results achieved from the full data sets.

Researchers were interested if fewer than the recommended 40 probes could be used – especially when placed in areas of high interest - but still obtain equivalent or better results. For this investigation, the computational results from a two-dimensional inlet and experimental results of an axisymmetric inlet were used. To find the areas of interest, a uniform sampling of all possible locations was run through a Monte Carlo simulation with a varying number of probes. A probability density function of the resultant distortion index was plotted. Certain probes are required to come within the desired accuracy level of the distortion index based on the full data set. For the experimental results, all three test cases could be characterized with 20 probes. For the axisymmetric inlet, placing 40 probes in select locations could get the results for parameters of interest within less than 10% of the exact solution for almost all cases. For the two dimensional inlet, the results were not as clear. 80 probes were required to get within 10% of the exact solution for all run numbers, although this is largely due to the small value of the exact result.

The sensitivity of each probe added to the experiment was analyzed. Instead of looking at the overall pattern established by optimizing probe placements, the focus is on varying the number of sampled probes from 20 to 40. The number of points falling within a 1\% tolerance band of the exact solution were counted as good points. The results were normalized for each data set and a general sensitivity function was found to determine the sensitivity of the results. A linear regression was used to generalize the results for all data sets used in this work. However, they can be used by directly comparing the number of good points obtained with various numbers of probes as well. The sensitivity in the results is higher when fewer probes are used and gradually tapers off near 40 probes. There is a bigger gain in good points when the number of probes is increased from 20 to 21 probes than from 39 to 40 probes.

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Conlon, Craig Fertig. "A study of ergonomic risk factors and interventions among aerospace engineers." Diss., Restricted to subscribing institutions, 2008. http://proquest.umi.com/pqdweb?did=1568065861&sid=1&Fmt=2&clientId=1564&RQT=309&VName=PQD.

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Reilly, Daniel. "An Investigation into Jet Engine Inlet Flow Characteristics for Turbine-Powered Helicopters." The Ohio State University, 2015. http://rave.ohiolink.edu/etdc/view?acc_num=osu1429608968.

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Ragozin, Konstantin. "Thrust Performance and Heat Load Modelling of Pulse Detonation Engines." Thesis, Luleå tekniska universitet, Institutionen för system- och rymdteknik, 2020. http://urn.kb.se/resolve?urn=urn:nbn:se:ltu:diva-82438.

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Pulse Detonation Engines (PDEs) are propulsion systems that use repeated detonations to generate thrust. Currently in early stages of development, PDEs have been theorised to have advantages over current deflagration based engines. Air-breathing PDEs could attain higher specific impulse values and operate at higher Mach numbers than today's air-breathing engines, while Pulse Detonation Rocket Engines (PDREs) could become a lighter, cheaper, and more reliable alternative to traditional rocket engines. There are still however, many technological hurdles to overcome before PDEs can be developed into practical propulsion systems, one major barrier being management of their immense heat loads. This thesis outlines the development of a numerical model for determining thrust performance and heat load characteristics of PDEs. The model is based on a set of analytical equations which characterise the gas dynamics inside the engine throughout it's cyclic process. Being numerically light -when compared to CFD analysis- the model allows for fast turnaround of results and the ability to sweep through parameters to determine optimum operating conditions to maximise engine performance and reduce heat load. In this study, the working principles of the model are described and it's outputs are validated against data from published experimental and numerical studies. The model is then used to conduct a comprehensive parametric study on the effects of various reactant combinations, operating conditions, and engine geometries on engine thrust, specific impulse and heat load. Lastly, a brief study is conducted on the feasibility of regenerative cooling for PDEs, using model outputs to determine if a heat balance can be achieved and the performance losses and complications that would result.

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Pakmehr, Mehrdad. "Towards verifiable adaptive control of gas turbine engines." Diss., Georgia Institute of Technology, 2013. http://hdl.handle.net/1853/49025.

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This dissertation investigates the problem of developing verifiable stable control architectures for gas turbine engines. First, a nonlinear physics-based dynamic model of a twin spool turboshaft engine which drives a variable pitch propeller is developed. In this model, the dynamics of the engine are defined to be the two spool speeds, and the two control inputs to the system are fuel flow rate and prop pitch angle. Experimental results are used to verify the dynamic model of JetCat SPT5 turboshaft engine. Based on the experimental data, performance maps of the engine components including propeller, high pressure compressor, high pressure, and low pressure turbines are constructed. The engine numerical model is implemented using Matlab. Second, a stable gain scheduled controller is described and developed for a gas turbine engine that drives a variable pitch propeller. A stability proof is developed for a gain scheduled closed-loop system using global linearization and linear matrix inequality (LMI) techniques. Using convex optimization tools, a single quadratic Lyapunov function is computed for multiple linearizations near equilibrium and non-equilibrium points of the nonlinear closed-loop system. This approach guarantees stability of the closed-loop gas turbine engine system. To verify the stability of the closed-loop system on-line, an optimization problem is proposed which is solvable using convex optimization tools. Through simulations, we show the developed gain scheduled controller is capable to regulate a turboshaft engine for large thrust commands in a stable fashion with proper tracking performance. Third, a gain scheduled model reference adaptive control (GS-MRAC) concept for multi-input multi-output (MIMO) nonlinear plants with constraints on the control inputs is developed and described. Specifically, adaptive state feedback for the output tracking control problem of MIMO nonlinear systems is studied. Gain scheduled reference model system is used for generating desired state trajectories, and the stability of this reference model is also analyzed using convex optimization tools. This approach guarantees stability of the closed-loop gain scheduled gas turbine engine system, which is used as a gain scheduled reference model. An adaptive state feedback control scheme is developed and its stability is proven, in addition to transient and steady-state performance guarantees. The resulting closed-loop system is shown to have ultimately bounded solutions with a priori adjustable bounded tracking error. The results are then extended to GS-MRAC with constraints on the magnitudes of multiple control inputs. Sufficient conditions for uniform boundedness of the closed-loop system is derived. A semi-global stability result is proven with respect to the level of saturation for open-loop unstable plants, while the stability result is shown to be global for open-loop stable plants. Simulations are performed for three different models of the turboshaft engine, including the nominal engine model and two models where the engine is degraded. Through simulations, we show the developed GS-MRAC architecture can be used for the tracking problem of degraded turboshaft engine for large thrust commands with guaranteed stability. Finally, a decentralized linear parameter dependent representation of the engine model is developed, suitable for decentralized control of the engine with core and fan/prop subsystems. Control theoretic concepts for decentralized gain scheduled model reference adaptive control (D-GS-MRAC) systems is developed. For each subsystem, a linear parameter dependent model is available and a common Lyapunov matrix can be computed using convex optimization tools. With this control architecture, the two subsystems of the engine (i.e., engine core and engine prop/fan) can be controlled with independent controllers for large throttle commands in a decentralized manner. Based on this D-GS-MRAC architecture, a "plug and play" (PnP) technology concept for gas turbine engine control systems is investigated, which allows us to match different engine cores with different engine fans/propellers. With this plug and play engine control architecture, engine cores and fans/props could be used with their on-board subordinate controllers ready for integration into a functional propulsion system. Simulation results for three different models of the engine, including the nominal engine model, the model with a new prop, and the model with a new engine core, illustrate the possibility of PnP technology development for gas turbine engine control systems.

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Ganine, Vladislav. "Model order reduction for prediction of turbine engine rotor vibration response in presence of parametric uncertainties." Thesis, McGill University, 2010. http://digitool.Library.McGill.CA:8881/R/?func=dbin-jump-full&object_id=92182.

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ALLGOOD, DANIEL CLAY. "AN EXPERIMENTAL AND COMPUTATIONAL STUDY OF PULSE DETONATION ENGINES." University of Cincinnati / OhioLINK, 2004. http://rave.ohiolink.edu/etdc/view?acc_num=ucin1095259010.

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Barone, Dominic L. "Investigation of TDLAS Measurements in a Scramjet Engine." University of Cincinnati / OhioLINK, 2010. http://rave.ohiolink.edu/etdc/view?acc_num=ucin1277130335.

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Andersson, Erik. "Preliminary design of a small-scale liquid-propellant rocket engine testing platform." Thesis, Luleå tekniska universitet, Rymdteknik, 2019. http://urn.kb.se/resolve?urn=urn:nbn:se:ltu:diva-77079.

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Propulsion system testing before mission operation is a fundamental requirement in any project. For both industrial and commercial entities within the space industry, complete system integration into a static test platform for functional and performance testing is an integral step in the system development process. Such a platform - if designed to be relatively safe, uncomplicated and reliable - can be an important tool within academia as well, giving researchers and students a possibility for practical learning and propulsion technology research. In this thesis, a preliminary design for a liquid-propellant rocket engine testing platform to be used primarily for academical purposes is developed. Included in the presented design is a bi-propellant Chemical Propulsion system, gas pressure fed with Gaseous Nitrogen and using Gaseous Oxygen as oxidiser and a 70 % concentrated ethanol-water mixture as fuel. The propellant assembly contains all necessary components for operating the system and performing combustion tests with it, including various types of valves, tanks and sensors. An estimation of the total preliminary cost of selected components is presented as well. Also part of the developed platform design is a small thrust chamber made of copper, water-cooled and theoretically capable of delivering 1000 N of thrust using the selected propellants. A list of operations to be performed before, during and after a complete combustion test is presented at the end of the document, together with a preliminary design of a Digital Control and Instrumentation System software. Due to time limitations, the software could not be implemented in a development program nor tested with simulated parameters as part of this thesis project.

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31

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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Potts, Ian. "Particle Redistribution in Serpentine Engine Inlets." The Ohio State University, 2020. http://rave.ohiolink.edu/etdc/view?acc_num=osu1595542100917769.

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Lawniczek, Baptiste. "Evolution of the methodology of weight estimation and engine feasibility in preliminary design." Thesis, KTH, Flygdynamik, 2019. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-261216.

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Som en del av Aerospace Engineering masterprogrammet som jag följde på KTH från augusti 2017 till juni 2018, fick jag slutligen en självständig studie i form av ett avhandlingsprojekt. Jag fick möjlighet att slutföra mitt 6-månaders examensprojekt på Safran Aircraft Engines i Villaroche, Frankrike. Safran är en internationell högteknologisk grupp som arbetar inom flygplanets framdrivning och utrustning, rymd- och försvarsmarknader. Som en del av de mekaniska aktiviteter som genomförs vid Produkt Innovation Lab är genomförandet av en för-dimensioneringsmetod för motorkonstruktioner nödvändiga för att få en uppfattning om för-dimensionering tidigare inom projektet och inom en rimlig tidsrymd. Huvudsyftet med detta projekt är att utveckla metoden som gör det möjligt att göra viktstatus och genomförbarhetsbedömningar av motorramkomponenterna i en preliminär konstruktionsfas och inom en begränsad tid. Detta papper fokuserar mer exakt på en äldre kommersiell motor mellanram. Denna metod måste leda till att en konsekvent modell skapas som ligger närmare kraven och specifikationerna. Slutsatsen av rapporten är att den implementerade pre-designmetoden möjliggör resultat med avseende på viktstatus och genomförbarhet som överensstämmer med specifikationerna. Vidare är beräkningstiden i linje med förväntningarna. Detta projekt har äntligen tillåtit att skapa en modell som kommer att tas som referens för att utveckla och designa nya motorer som har en konfiguration som liknar äldre kommersiella motorer.
This paper aims to develop and to validate a methodology to realize pre-sizing studies on aircraft engine structural frames for Safran Aircraft Engines Product Innovation Lab activities. The members of this team are in charge of creating new propulsion systems architectures in accordance with product strategy guidelines or airframer needs. Aerospace industry being highly competitive, the Product Innovation Lab must be able to respond quickly and precisely to any demand emerging from aircraft manufacturers or strategy team.The main purpose of this project is to improve the methodology permitting to make weight status and feasibility estimations of the engine frame components in a preliminary design phase and in a limited amount of time. This methodology must lead to the creation of a consistent model that is closer to the requirements and specifications imposed. This paper more precisely focus on legacy commercial engine structural frame. Reflection has been conducted on the creation of a simplified parametrized model of an existing commercial engine structural frame and on the way to mesh it in order to find a good compromise between results fidelity and computation time. Regarding the weight status and feasibility results obtained with a first model, an optimization of the model configuration has finally been conducted in order to get results that fit with the specifications.Conclusion of the report is that the pre-sizing methodology can be adapted to existing commercial engine structural frame configuration. Results obtained in terms of weight status and feasibility are in accordance with the specifications and the computation time is in agreement with the expectations. It has permitted to create a model that will be taken as a reference to develop and design new engines having a configuration similar to the legacy commercial engine considered in this study. For that purpose, iterations and optimizations will be conducted on the simplified model implemented during the project in order to determine a new configuration of the pre-sized intermediate frame model which sticks with reality i.e. that respects the provided feasibility specifications.Note that due to the public nature of this report, sensitive information and data used and obtained during the project have been removed from the present paper. Nevertheless, the methodology followed has been presented and discussed in detail. Relative deviations between the results obtained and reference values have also been exposed in order to give the reader an idea of implemented model consistency.

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34

Giuliani, James Edward. "Jet Engine Fan Response to Inlet Distortions Generated by Ingesting Boundary Layer Flow." The Ohio State University, 2016. http://rave.ohiolink.edu/etdc/view?acc_num=osu1468564279.

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35

Kumar, Abhinav. "Flow control optimization in a jet engine serpentine inlet duct." [College Station, Tex. : Texas A&M University, 2007. http://hdl.handle.net/1969.1/ETD-TAMU-1399.

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North, Gary S. "Metal Coupon Testing in an Axial Rotating Detonation Engine for Wear Characterization." Wright State University / OhioLINK, 2020. http://rave.ohiolink.edu/etdc/view?acc_num=wright1588770787704665.

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Driscoll, Robert B. "Investigation of Sustained Detonation Devices: the Pulse Detonation Engine-Crossover System and the Rotating Detonation Engine System." University of Cincinnati / OhioLINK, 2016. http://rave.ohiolink.edu/etdc/view?acc_num=ucin1459155478.

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38

Yuan, Zheng Shan. "Oscillatory flow and heat transfer in a Stirling engine regenerator." Case Western Reserve University School of Graduate Studies / OhioLINK, 1993. http://rave.ohiolink.edu/etdc/view?acc_num=case1056988063.

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39

Lundmark, Martin. "Numerical study with computational fluid dynamics of hybrid rocket engine." Thesis, Luleå tekniska universitet, Rymdteknik, 2020. http://urn.kb.se/resolve?urn=urn:nbn:se:ltu:diva-82037.

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In this thesis a Large Eddy Simulation (LES) of a hybrid rocket engine burning ethylene (C2H4) in nitrous oxide (N2O) is explored. This is done primarily using a solver and solution scheme provided by the Swedish Defence Research Agency (FOI) and an (at this date) unpublished chemistry model. This sheds light on some transiet behaviour of a prior experiment conducted with a model engine that the simulation was based on. Due to time constraints the simulation did not cover the full test of the engine. The results confirm predictions from the experiment that the propellant was fuel rich. Some insight on how oxidizer swirl propagates throughout the engine was discovered as well.

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Suchocki, James Alexander. "Operational Space and Characterization of a Rotating Detonation Engine Using Hydrogen and Air." The Ohio State University, 2012. http://rave.ohiolink.edu/etdc/view?acc_num=osu1330266587.

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41

Driscoll, Robert B. "EXPERIMENTAL INVESTIGATION OF SHOCK TRANSFER AND SHOCK INITIATED DETONATION IN A DUAL PULSE DETONATION ENGINE CROSSOVER SYSTEM." University of Cincinnati / OhioLINK, 2013. http://rave.ohiolink.edu/etdc/view?acc_num=ucin1378113995.

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Curtner, Charles R. "Output Feedback with Output Tracking, with Application to a Turbofan Engine." University of Cincinnati / OhioLINK, 2014. http://rave.ohiolink.edu/etdc/view?acc_num=ucin1408709619.

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43

Eastbourn, Scott Michael. "Modeling and Simulation of a Dynamic Turbofan Engine Using MATLAB/Simulink." Wright State University / OhioLINK, 2012. http://rave.ohiolink.edu/etdc/view?acc_num=wright1340582603.

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Allenstein, Jacob T. "An Investigation of Jet Engine Test Cell Exhaust Stack Aerodynamics and Performance through Scale Model Test Studies and Computational Fluid Dynamics Results." The Ohio State University, 2020. http://rave.ohiolink.edu/etdc/view?acc_num=osu1586515794023938.

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DRENSKY, GEORGE K. "EXPERIMENTAL INVESTIGATION OF COMPOSITE MATERIAL EROSION CHARACTERISTICS UNDER CONDITIONS ENCOUNTERED IN TURBOFAN ENGINES." University of Cincinnati / OhioLINK, 2007. http://rave.ohiolink.edu/etdc/view?acc_num=ucin1178118863.

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46

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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47

KAMARAJ, JAYACHANDRAN. "MODELING AND SIMULATION OF SINGLE SPOOL JET ENGINE." University of Cincinnati / OhioLINK, 2004. http://rave.ohiolink.edu/etdc/view?acc_num=ucin1073935505.

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Bond, Ryan Bomar. "Reynolds-Averaged Navier-Stokes Analysis of the Flow through a Model Rocket-Based Combined Cycle Engine with an Independently-Fueled Ramjet Stream." NCSU, 2003. http://www.lib.ncsu.edu/theses/available/etd-08132003-171258/.

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A new concept for the low speed propulsion mode in rocket based combined cycle (RBCC) engines has been developed as part of the NASA GTX program. This concept, called the independent ramjet stream (IRS) cycle, is a variation of the traditional ejector ramjet (ER) design and involves the injection of hydrogen fuel directly into the air stream, where it is ignited by the rocket plume. Experiments and computational fluid dynamics (CFD) are currently being used to evaluate the feasibility of the new design. In this work, a Navier-Stokes code valid for general reactive flows is applied to the model engine under cold flow, ejector ramjet, and IRS cycle operation. Pressure distributions corresponding to cold-flow and ejector ramjet operation are compared with experimental data. The engine response under independent ramjet stream cycle operation is examined for different reaction models and grid sizes. The engine response to variations in fuel injection is also examined. Mode transition simulations are also analyzed both with and without a nitrogen purge of the rocket. The solutions exhibit a high sensitivity to both grid resolution and reaction mechanism, but they do indicate that thermal throat ramjet operation is possible through the injection and burning of additional fuel into the air stream. The solutions also indicate that variations in fuel injection location can affect the position of the thermal throat. The numerical simulations predicted successful mode transition both with and without a nitrogen purge of the rocket; however, the reliability of the mode transition results cannot be established without experimental data to validate the reaction mechanism.

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Lindkvist, Oskar. "Model Adaptation of a Mixed Flow Turbofan Engine." Thesis, Luleå tekniska universitet, Institutionen för system- och rymdteknik, 2020. http://urn.kb.se/resolve?urn=urn:nbn:se:ltu:diva-80667.

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Gas turbine performance models are usually created in an object oriented manner, where different standard components are connected to form the complete model. The characteristics of these components are often represented by component maps and empirical correlations. However, engine specific component characteristics are seldom available to anyone outside of the manufacturers. It is therefore very common for researchers to use publicly accessible or generic component maps instead. But in order to reduce prediction errors the maps have to be modified to fit any specific engine. This thesis work investigates the process of adapting a parametric turbofan engine model to a limited amount of test-data using the propulsion program EVA. Steady state test-data was generated using an initial reference model with SLS operating conditions. Another engine model with different fan, compressor and turbine maps was then used in the adaptation. An initial on-design model was adapted to the highest power test-data point. This model is based on aerothermodynamic equations and is used as a reference to scale the generic component maps to. A sensitivity analysis was done at this point in order to find dependencies between unknown component parameters and test data. These were then included in the cycle solver which employs a version of the Newton-Raphson method. After the fan and compressor maps had been scaled to the design point they were adapted to test-data by adjusting the mass flow parameters in a direct search optimizer. Finally, speed lines in the fan and compressor maps were relabeled to reduce rotor speed errors. The adapted performance model was then validated against the reference model at a few flying conditions. The performance model results demonstrate that it is possible to greatly reduce prediction errors by only adjusting the corrected mass flow in fan and compressor maps. Additionally, rotor speed errors could successfully be corrected as a final step in the adaptation by relabeling speed lines in the component maps. When validated, the adapted model had a maximum parameter error of 1.5%.

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Shekhar, Anjali. "Study and Numerical Simulation of Unconventional Engine Technology." University of Cincinnati / OhioLINK, 2018. http://rave.ohiolink.edu/etdc/view?acc_num=ucin1544098607675961.

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Dissertations / Theses on the topic 'Aerospike engines'

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