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1
Hackaday, Gary L. "Thrust augmentation for a small turbojet engine." Thesis, Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 1999. http://handle.dtic.mil/100.2/ADA362981.
Full textAbstract:
Thesis (M.S. in Aeronautical Engineering) Naval Postgraduate School, March 1999.<br>Thesis advisor(s): Garth V. Hobson. "March 1999". Includes bibliographical references (p. 75). Also available online.
2
MEDICI, GIOVANNI. "Thrust vectoring on an UCAV airplane (thrust vectoring engine model): advanced control system." Doctoral thesis, Politecnico di Torino, 2013. http://hdl.handle.net/11583/2508286.
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Since the early development of the Sky-X UAV demonstra- tor, Alenia Aermacchi has been studying a thrust vectored version of that airplane. For this reason, in 2010, Alenia Aermacchi offered a three years PhD funding to the Depart- ment of Aerospace Engineering of Politecnico di Torino.
In particular a vane based thrust vectoring system had to be studied, the development of open and flexible software tools was a priority in the project. The tools had to be cus- tomizable and a general estimation model of the system had to be developed.
The research analysis took place during three years and involved several areas of the company and of aerospace en- gineering.
In particular three main areas were covered in the study: propulsion, aerodynamics , flight dynamics.
3
Yarlagadda, Santosh. "Performance Analysis of J85 Turbojet Engine Matching Thrust with Reduced Inlet Pressure to the Compressor." University of Toledo / OhioLINK, 2010. http://rave.ohiolink.edu/etdc/view?acc_num=toledo1271367584.
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4
Bartosh, Brady J. "Thrust measurement of a split-path, valveless pulse detonation engine." Thesis, Monterey, Calif. : Naval Postgraduate School, 2007. http://bosun.nps.edu/uhtbin/hyperion-image.exe/07Dec%5FBartosh.pdf.
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Thesis (M.S. in Astronautical Engineering)--Naval Postgraduate School, December 2007.<br>Thesis Advisor(s): Brophy, Christopher M. "December 2007." Description based on title screen as viewed on January 17, 2008. Includes bibliographical references (p. 95-96). Also available in print.
5
Boggs, George Lemuel IV. "Turbine Engine Thrust Measurements Using a Non-Intrusive Acoustic Technique." Thesis, Virginia Tech, 2019. http://hdl.handle.net/10919/90299.
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Experiments were conducted to measure the thrust generated from a commercial jet engine. This thrust estimation was done using a pneumatic horn as the sound source with two arrays of microphones directly across the exhaust stream. The two arrays were separated by an axial distance downstream. Exhaust centerline measurements were taken at varying engine conditions, specifically; 30%, 50%, 60%, 70%, 80% and 100% engine power. An acoustic thrust estimation showed good agreement with measured thrust during the test campaign. In addition, a full traverse of the acoustic rig through the exhaust stream was completed for the purpose of tomography reconstruction. This reconstruction technique was able to pick up key features of the flow field.<br>Master of Science
6
Schaefermeyer, M. Ryan. "Aerodynamic Thrust Vectoring For Attitude Control Of A Vertically Thrusting Jet Engine." DigitalCommons@USU, 2011. https://digitalcommons.usu.edu/etd/1237.
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NASA’s long range vision for space exploration includes human and robotic missions to extraterrestrial bodies including the moon, asteroids and the martian surface. All feasible extraterrestrial landing sites in the solar system are smaller and have gravitational fields of lesser strength than Earth’s gravity field. Thus, a need exists for evaluating autonomous and human-piloted landing techniques in these reduced-gravity situations. A small-scale, free-flying, reduced-gravity simulation vehicle was designed by a group of senior mechanical engineering students with the help of faculty and graduate student advisors at Utah State University during the 2009-2010 academic year. The design reproduces many of the capabilities of NASA’s 1960s era lunar landing research vehicle using small, inexpensive modern digital avionics instead of the large, expensive analog technology available at that time. The final vehicle design consists of an outer maneuvering platform and an inner gravity offset platform. The two platforms are connected through a set of concentric gimbals which allow them to move in tandem through lateral, vertical, and yawing motions, while remaining independent of each other in rolling and pitching motions. A small radio-controlled jet engine was used on the inner platform to offset a fraction of Earth’s gravity (5/6th for lunar simulations), allowing the outer platform to act as though it is flying in a reduced-gravity environment. Imperative to the stability of the vehicle and fidelity of the simulation, the jet engine must remain in a vertical orientation to not contribute to lateral motions. To this end, a thrust vectoring mechanism was designed and built that, together with a suite of sensors and a closed loop control algorithm, enables precise orientation control of the jet engine. Detailed designs for the thrust vectoring mechanisms and control avionics are presented. The thrust vectoring mechanism uses thin airfoils, mounted directly behind the nozzle, to deflect the engine’s exhaust plume. Both pitch and yaw control can be generated. The thrust vectoring airfoil sections were sized using the two-dimensional airfoil section compressible-flow CFD code, XFOIL, developed at the Massachusetts Institute of Technology. Because of the high exhaust temperatures of the nozzle plume, viscous calculations derived from XFOIL were considered to be inaccurate. XFOIL was run in inviscid flow mode and viscosity adjustments were calculated using a Utah State University-developed compressible skin friction code. A series of ground tests were conducted to demonstrate the thrust vectoring system’s ability to control the orientation of the jet engine. Detailed test results are presented.
7
Dunn, Francis X. "Thrust modulated multivariable control of the GE21 engine using the LQG/LTR method." Thesis, Massachusetts Institute of Technology, 1986. http://hdl.handle.net/1721.1/15000.
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Thesis (M.S.)--Massachusetts Institute of Technology, Dept. of Electrical Engineering and Computer Science, 1986.<br>MICROFICHE COPY AVAILABLE IN ARCHIVES AND ENGINEERING<br>Bibliography: leaves 173-175.<br>by Francis X. Dunn.<br>M.S.
8
Alves, Wilton Fernandes. "Development of experimental firing test stand to study the rocket engine thrust characteristics." Instituto Tecnológico de Aeronáutica, 2008. http://www.bd.bibl.ita.br/tde_busca/arquivo.php?codArquivo=2324.
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The main aim of this work is to present the specification of an experimental firing test stand of liquid rocket engine (LRE), comprising the main design, the instrumentation of measurement system, the data acquisition system, the operating manual, as well as the methodology to perform laboratory work for determination of a LRE thrust characteristics in atmospheric conditions. Initially it is presented a theoretical basement of LRE in general and concerning the laboratory work. After that it is proposed a methodology for execution of laboratory work using resources of information technology, which will allow the automatic and remote functioning of the test stand, and it will give to the users the inputs necessaries to realization of tests and attainment of reliable results. The specification of the test stand is result of calculations implemented in MathCAD program in way of algorithms presented in appendix of this work. The control of mass flow rates of propellant by automatic pressure regulators and valves, as well as the data acquisition of test stand is carried out by Labview program in a NI PXI platform. The instrumentation of measurement system will make possible online measurements of temperatures, pressures, mass flow rates and thrust force related to the tests. It is presented also a preliminary analysis of type B uncertainties of test stand system, and a comparative analysis between designed LRE with similar rocket engine of a test stand in operation.
9
MacLeod, James D. (James Donald) Carleton University Dissertation Engineering Aeronautical. "A derivation of gross thrust for a sea-level jet engine test cell." Ottawa, 1988.
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10
Gullia, Alessandro. "Thrust and Flow Prediction in Gas Turbine Engine Indoor Sea-Level Test Cell Facilities." Thesis, Cranfield University, 2006. http://dspace.lib.cranfield.ac.uk/handle/1826/7496.
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The principal aim of this research was to provide a detailed understanding of the
performance of gas turbine engines inside indoor sea-level test beds. In particular the
evaluation of both thrust correction factors and the estimation of the mass flow
entering the test cell were at the core of the research.
The project has been fully sponsored by Rolls-Royce pIc. Initially, their principal
objective was to assess the relevance and accuracy of CFD when applied to thrust
measurement inside indoor test beds with an intended outcome of minimising the use
of expensive experimental measurements.
The different system interfaces and accounting systems for in-flight conditions,
available in the open literature have been developed and adapted for indoor
environments. This has led to the definition of three different thrust correction
equations using alternative definitions of thrust correction factor. Aero-dynamic
principles have been applied for the derivation of one-dimensional relationships for
the calculation of each thrust correction factor using generic engine-cell performance
and dimensions.
A one-dimensional analytical model has been developed to represent the enginedetuner
ejector pump. This is able to characterise the engine-cell system performance
and is used as the main tool for providing a matching procedure capable of predicting
the cell entrainment ratio.
By processing experimental data relevant to different engine-cell configurations
through the ejector pump analytical model, a method for achieving the entrainment
ratio control inside the cell has been identified.
The CFD work has been concentrated into three main activities:
• A quantitative extrapolation of the thrust correction factors including, the
pre-entry force, the external and the total bellmouth force, the throat stream
force, the intake momentum drag and the base drag.
• The representation of the engine-detuner ejector performance for a
variety of engine-cell configurations.
• The modelling of the generic test cell components including the inlet
stack, the cascade elbow, the exhaust stack & the blast basket.
The outcomes of this research have been very successful in enhancing the validity of
the thrust correction equations developed .. In particular, the use of a one-dimensional
approach in their estimation has been shown to be fully justified. The work has also
emphasised the value of CFD in supporting the derivation of the matching procedure
for predicting and controlling cell entrainment ratio. Indeed, one of the strongest
outcomes of this work has been the conclusion that both the engine-cell characteristic
lines computed with the one-dimensional model and those computed with CFD for
different cell configurations are almost identical.
In addition, the use of CFD as a tool for the quantitative evaluation of the thrust
correction factors has been established. Finally, the CFD results have facilitated an
enhanced understanding of the complex flow structure inside indoor test cells
11
Wu, Fuh-Eau. "Aero engine life evaluated for combined creep and fatigue, and extended by trading-off excess thrust." Thesis, Cranfield University, 1994. http://dspace.lib.cranfield.ac.uk/handle/1826/10517.
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This thesis investigates the concept of thrust rating as a means towards reducing the life cycle costs of engine ownership. Towards this end, this thesis has discussed the concept of thrust rating, developed computer programs for mechanical load type failures, which include creep, LCF, and combinations thereof, and conducted simulations of improving life usage and reducing life cycle costs. A study was performed on a military engine, under an original design mission mix, that showed significant gains in creep-LCF life of the HPT blade could be achieved, especially With the recently proposed and presumably more accurate criterion- ductility exhaustion, by thrust rating. The savings were expressed in terms of an approximate reduced life accumulation rates and life cycle costs. The net result was a 50% increase in creep-LCF life with a savings of $50.4 million. These calculations were based on a Feet of 300 engines having the designed lifetime of 8,000 operating hours per engine. Throughout the thesis, mention is also made of employing the thrust rating concept on other engines. To this end, the thesis will also give a blueprint for conducting a feasibility study to employ thrust rating as a maintenance tool. In addition to the technical aspects, the role of maintenance and aircraft operations policy will also be studied to determine the interrelationships that exist between thrust rating technology and its practical application.
12
Sarwade, Rohit Foster Winfred A. "Life prediction analysis of a subscale rocket engine combustor using a fluid-thermal-structural model." Auburn, Ala., 2006. http://repo.lib.auburn.edu/2006%20Spring/master's/SARWADE_ROHIT_49.pdf.
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13
Allenstein, Jacob T. "An Investigation in Gold-Plating Scaled Turbofan Engine Simulators through Means of Aerodynamic and Load Cell Thrust Measurements with Comparisons to Full-Scale Engine Results." The Ohio State University, 2013. http://rave.ohiolink.edu/etdc/view?acc_num=osu1386061117.
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14
Wu, F.-E. "Aero engine life evaluated for combined creep and fatigue, and extended by trading-off excess thrust." Thesis, Cranfield University, 1994. http://dspace.lib.cranfield.ac.uk/handle/1826/10517.
Full textAbstract:
This thesis investigates the concept of thrust rating as a means
towards
reducing the life cycle costs of engine ownership. Towards this end,
this
thesis has discussed the concept of thrust rating, developed computer
programs for mechanical load type failures, which include creep, LCF, and
combinations thereof, and conducted simulations of improving life usage and
reducing life cycle costs.
A
study was performed on a military engine, under an original design
mission mix, that showed significant gains in creep-LCF life of the HPT blade
could be achieved, especially With the recently proposed and presumably more
accurate criterion- ductility exhaustion, by thrust rating. The savings were
expressed in terms of an approximate reduced life accumulation rates and life
cycle costs. The net result was a 50% increase in creep-LCF life with a
savings of $ 50.4 million. These calculations were based on a Feet of 300
engines having the designed lifetime of 8,000 operating hours per engine.
Throughout the thesis, mention is also made of employing the thrust
rating concept on other engines. To this end, the thesis will also give a
blueprint for conducting a feasibility study to employ thrust rating as a
maintenance tool. In addition to the technical aspects, the role of maintenance
and aircraft
operations policy will also be studied to determine the
interrelationships that exist between thrust rating technology and its
practical application.
15
Wittmers, Nicole K. "Direct-connect performance evaluation of a valveless pulse detonation engine." Thesis, Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 2004. http://library.nps.navy.mil/uhtbin/hyperion/04Dec%5FWittmers.pdf.
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16
Sakai, Tadashi. "A Study of Variable Thrust, Variable Specific Impulse Trajectories for Solar System Exploration." Diss., Georgia Institute of Technology, 2004. http://hdl.handle.net/1853/4904.
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A study has been performed to determine the advantages and disadvantages of variable thrust and variable specific impulse (Isp) trajectories for solar system exploration.
There have been several numerical research efforts for variable thrust, variable Isp, power-limited trajectory optimization problems. All of these results conclude that variable thrust, variable Isp (variable specific impulse, or VSI) engines are superior to constant thrust, constant Isp (constant specific impulse, or CSI) engines. However, most of these research efforts assume a mission from Earth to Mars, and some of them further assume that these planets are circular and coplanar. Hence they still lack the generality.
This research has been conducted to answer the following questions:
- Is a VSI engine always better than a CSI engine or a high thrust engine for any mission to any planet with any time of flight considering lower propellant mass as the sole criterion?
- If a planetary swing-by is used for a VSI trajectory, is the fuel savings of a VSI swing-by trajectory better than that of a CSI swing-by or high thrust swing-by trajectory?
To support this research, an unique, new computer-based interplanetary trajectory calculation program has been created. This program utilizes a calculus of variations algorithm to perform overall optimization of thrust, Isp, and thrust vector direction along a trajectory that minimizes fuel consumption for interplanetary travel. It is assumed that the propulsion system is power-limited, and thus the compromise between thrust and Isp is a variable to be optimized along the flight path. This program is capable of optimizing not only variable thrust trajectories but also constant thrust trajectories in 3-D space using a planetary ephemeris database. It is also capable of conducting planetary swing-bys.
Using this program, various Earth-originating trajectories have been investigated and the optimized results have been compared to traditional CSI and high thrust trajectory solutions. Results show that VSI rocket engines reduce fuel requirements for any mission compared to CSI rocket engines. Fuel can be saved by applying swing-by maneuvers for VSI engines, but the effects of swing-bys due to VSI engines are smaller than that of CSI or high thrust engines.
17
Micklewright, Alan J. "An analysis of single-engine rate-of-climb capabilities and thrust requirements of the S-3 and ES-3 aircraft in support of the TF34 Engine Component Improvement Program." Thesis, Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 1995. http://handle.dtic.mil/100.2/ADA302899.
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18
Martinez, Anna. "Design and manufacturing of a thrust measurement system for a micro jet engine : Enabling in-flight drag estimation for subscale aircraft testing." Thesis, Linköpings universitet, Fluida och mekatroniska system, 2018. http://urn.kb.se/resolve?urn=urn:nbn:se:liu:diva-149288.
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Good estimation of aerodynamic coefficients is of fundamental importance in the design and development process of an aircraft. Generally, these parameters are obtained using analytical, numerical and experimental methods, which are sometimes either inaccurate or very expensive. The use of subscale aircraft is becoming increasingly common in the study and evaluation of new aircraft concepts. Flight testing results in an efficient solution for obtaining parameters that can define drag characteristics. This project presents a solution for achieving the drag aerodynamic model from the design and manufacturing of a micro engine thrust measuring system integrated on subscale aircraft. Strain gauge technology permits to identify the stresses that the engine forces cause to the aircraft internal structure by analysing the strain of several strategic zones of the engine mounting created for this purpose. Different structural support geometries have been presented and stress-analysed together with the design of the appropriate strain gauge model conguration in order to select and manufacture a system that represents a good compromise between all the requirements while ensuring the quality and accuracy of the data acquired. After calibration, installation and set-up, the system is ready for real in-flight measurements.
19
Oliver, Michael James. "A STUDY ON THE PHYSICS OF ICE ACCRETION IN A TURBOFAN ENGINE ENVIRONMENT." Case Western Reserve University School of Graduate Studies / OhioLINK, 2013. http://rave.ohiolink.edu/etdc/view?acc_num=case1363875844.
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20
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.
21
Mutschall, Marcel. "Die Genauigkeit einer vereinfachten Berechnung der Steigzeit von Flugzeugen." Aircraft Design and Systems Group (AERO), Department of Automotive and Aeronautical Engineering, Hamburg University of Applied Sciences, 2018. http://d-nb.info/1175497711.
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Ziel - Die Zeit die ein Flugzeug benötigt, um auf eine bestimmte Höhe zu steigen (die
Steigzeit) kann mit einer Formel berechnet werden, die vereinfachend annimmt, dass die
Steiggeschwindigkeit über dem gesamten Steigflug mit zunehmender Höhe linear abnimmt.
Ziel der Untersuchung ist, zu ermitteln, ob die Annahme einer linear abnehmenden
Steiggeschwindigkeit realistisch ist bzw. welche Fehler sich aus der Annahme ergeben.
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Methode - Mit der Höhe ändern sich Parameter wie Luftdichte, Widerstand, Schub und damit
auch die optimale Fluggeschwindigkeit für den Steigflug. Die Parameter beeinflussen sich
dabei gegenseitig. Der Schub wird dabei nach drei unterschiedlichen Methoden berechnet,
gegeben von Bräunling, Scholz und Howe. Analysiert wird der Verlauf des Schubes mit der
Höhe und der Verlauf der Steiggeschwindigkeit mit der Höhe für jede der drei
Schubberechnungen. Abschließend wird für jede Schubberechnung die Steigzeit verglichen
wie sie sich ergibt a) aus der einfachen Formel und b) aus einer Integrationsberechnung, bei
der der Verlauf der Steiggeschwindigkeit durch eine Funktion beschrieben wird.
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Ergebnisse - Die drei Schubberechnungen liefern ausgehend vom gleichen Startschub
unterschiedliche Schübe in der Höhe. In die Methode nach Bräunling gehen mehr Parameter
ein als in die anderen beiden Methoden. Es kann angenommen werden, dass die Methode
nach Bräunling genauer ist, der Beweis kann aber nicht geführt werden. Der Schub nach
Scholz und Howe fällt nahezu linear mit der Höhe ab. Der Schubverlauf nach Bräunling zeigt
eine deutliche Nichtlinearität. Es wird die Steigzeit von 0 km auf 11 km Höhe berechnet nach
a) und b), mit jeder der drei Schubberechnungen. Dabei wird jeweils der Unterschied in der
Steigzeit ermittelt. Aufgrund der Nichtlinearität im Schubverlauf zeigt die Methode nach
Bräunling dann auch den größten Unterschied zwischen den Berechnungsmethoden von
7,1 %. Bei einer Schubberechnung nach Scholz ergeben sich 1,7 % und nach Howe 1,4 %.
Wenn bereits zu Beginn Vereinfachungen, z.B. bezüglich des Triebwerksschubes,
vorgenommen wurden, ist es in Hinblick auf den Aufwand und die zu erreicheneden
Ergebnisse möglich, und zum Teil sinnvoll, die Berechnungen der Steigzeit mittels linearer
Abnahme der vertikalen Geschwindigkeit durchzuführen. Es wird ausdrücklich darauf
hingewiesen, dass es hier um den Vergleich von zwei Methoden zur Berechnung der Steigzeit
geht und nicht um die Bewertung von Methoden zur Schubberechnung (für die keine
Vergleichswerte vorlagen).
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Praktischer Nutzen - Es konnte festgestellt werden, dass eine einfache Formel zur
Berechnung der Steigzeit mit geringem Fehler angewandt werden kann - insbesondere wenn
Methoden zur Schubberechnung vorliegen, bei denen der Schub annähernd linear mit der
Höhe abnimmt. Bei großem Aufwand und realitätsnaher Betrachtung, z.B. nach Bräunling,
führt der lineare Ansatz jedoch zu einem zu großen Fehler. Hierfür sollte die Berechnung der
Steigzeit mittels Integration durchgeführt werden.
22
Harl, Nathan Robert. "Low-thrust control of a lunar orbiter." Diss., Rolla, Mo. : University of Missouri-Rolla, 2007. http://scholarsmine.mst.edu/thesis/pdf/Thesis_09007dcc80318673.pdf.
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Thesis (M.S.)--University of Missouri--Rolla, 2007.<br>Vita. The entire thesis text is included in file. Title from title screen of thesis/dissertation PDF file (viewed October 24, 2007) Includes bibliographical references (p. 95-98).
23
Bensel, Artur. "Characteristics of the Specific Fuel Consumption for Jet Engines." Aircraft Design and Systems Group (AERO), Department of Automotive and Aeronautical Engineering, Hamburg University of Applied Sciences, 2018. http://d-nb.info/1175791237.
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Purpose of this project is a) the evaluation of the Thrust Specific Fuel Consumption (TSFC) of jet engines in cruise as a function of flight altitude, speed and thrust and b) the determination of the optimum cruise speed for maximum range of jet airplanes based on TSFC characteristics from a). Related to a) a literature review shows different models for the influence of altitude and speed on TSFC. A simple model describing the influence of thrust on TSFC seems not to exist in the literature. Here, openly available data was collected and evaluated. TSFC versus thrust is described by the so-called bucket curve with lowest TSFC at the bucket point at a certain thrust setting. A new simple equation was devised approximating the influence of thrust on TSFC. It was found that the influence of thrust as well as of altitude on TSFC is small and can be neglected in cruise conditions in many cases. However, TSFC is roughly a linear function of speed. This follows already from first principles. Related to b) it was found that the academically taught optimum flight speed (1.316 times minimum drag speed) for maximum range of jet airplanes is inaccurate, because the derivation is based on the unrealistic assumption of TSFC being constant with speed. Taking account of the influence of speed on TSFC and on drag, the optimum flight speed is only about 1.05 to 1.11 the minimum drag speed depending on aircraft weight. The amount of actual engine data was extremely limited in this project and the results will, therefore, only be as accurate as the input data. Results may only have a limited universal validity, because only four jet engine types were analyzed. One of the project's original value is the new simple polynomial function to estimate variations in TSFC from variations in thrust while maintaining constant speed and altitude.
24
Weber, Fabian. "Optical Analysis of the Hydrogen Cooling Film in High Pressure Combustion Chambers." Thesis, Luleå tekniska universitet, Rymdteknik, 2019. http://urn.kb.se/resolve?urn=urn:nbn:se:ltu:diva-76872.
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For performance optimisation of modern liquid cryogenic bipropellant rocket combustion chambers, one component which plays an important role in reducing the wall side heat flux, is the behaviour of the cooling film. At the Institute of Space Propulsion of the German Aerospace Center (DLR) in Lampoldshausen, hot test runs have been performed using the experimental combustion chamber BKM, to investigate the wall side heat flux which is -- among other factors -- dependent on cooling film properties. To gain more insight into the film behaviour under real rocket-like conditions, optical diagnostics have been applied. The chosen methods were shadowgraphy and OH* imaging producing optical data sets which are analysed in this study. In this context, a description of the necessary background information is given, concerning rocket combustion chambers, film cooling and optical diagnostics of O2/H2 combustion. The applied methodology for optical analysis is described, followed by a presentation of the results. During the test campaign, it became clear that the optical setup was not optimised for creating meaningful shadowgraphy recordings which is why the shadowgraphy data has to be treated as flame emission imaging. The behaviour of the gas layer adjacent to the chamber wall could be characterised based on qualitative (luminosity, LOx shadow, reflection, recirculation zone and flame shape) and quantitative (layer thickness, layer length, pressure conditions) analysis. The thickness could be identified for each load step and an average length of the layer was found as well. OH* imaging has been used supplementary to support the observations from the flame emission images. An in depth frame by frame analysis was not possible due to time constraints. However, the time averaged images yielded results in accordance to the flame emission and could give a relative figure for the temperature distribution in the combustion volume. An artefact in the data was found, stemming presumably from the image intensifier. This artefact needs to be researched for a future error reduction in the data of this and other campaigns. Additionally, the thickness of the layer suggested a correlation to the models for film cooling efficiency. Such a correlation could not be established. Nevertheless, the film cooling models show the same behaviour as the data obtained from the flame emission imaging. Finally, suggestions are given how the data analysis and the optical setup could be improved for future, similar campaigns.
25
Dwyer, Michael E. "Quasi-optimal steady state and transient maneuvers with and without thrust vectoring." Thesis, This resource online, 1993. http://scholar.lib.vt.edu/theses/available/etd-09292009-020230/.
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26
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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Josselyn, Scott B. "Optimization of low thrust trajectories with terminal aerocapture." Thesis, Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 2003. http://library.nps.navy.mil/uhtbin/hyperion-image/03Jun%5FJosselyn.pdf.
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Thesis (Aeronautical and Astronautical Engineer)--Naval Postgraduate School, June 2003.<br>Thesis advisor(s): I. Michael Ross, Steve Matousek. Includes bibliographical references (p. 149-150). Also available online.
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Rust, Jack W. "Fuel optimal low thrust trajectories for an asteroid sample return mission." Thesis, Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 2005. http://library.nps.navy.mil/uhtbin/hyperion/05Mar%5FRust.pdf.
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Miller, Robert E. Hartfield Roy J. "Design and testing of a gas distribution method for pulsed inductive thruster." Auburn, Ala, 2008. http://hdl.handle.net/10415/1405.
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Ho, Ivan Chin Kian. "Investigation on novel methods to increase specific thrust in pulse detonation engines via imploding detonations." Thesis, Monterey, California : Naval Postgraduate School, 2009. http://edocs.nps.edu/npspubs/scholarly/theses/2009/Dec/09Dec%5FHo%5FIvan.pdf.
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Thesis (M.S. in Mechanical Engineering and M.S. in Applied Physics)--Naval Postgraduate School, December 2009.<br>Thesis Advisor(s): Sinibaldi, Jose. O. ; Brophy, Christopher M. "December 2009." Description based on title screen as viewed on January 27, 2010. Author(s) subject terms: Pulse Detonation Engines, Shock Wave, Detonation Implosions, Induction Length, Detonation Cell Width, Deflagration-to-Detonation Transition, Specific Thrust. Includes bibliographical references (p. 93-94). Also available in print.
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Tsiotras, Panagiotis. "Goddard-problem variants." Thesis, Virginia Polytechnic Institute and State University, 1987. http://hdl.handle.net/10919/80041.
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The problem of maximizing the altitude of a rocket in vertical flight has been extensively analyzed by many writers since the early days of rocketry. In the beginning, solutions were obtained using the classical theory of the Calculus of Variations, and later using Optimal Control theory. For strict assumptions on the drag law and the thrust, solutions were found, even in a closed, analytic form. Nevertheless, for more realistic conditions, the problem becomes a very complex one, and the solution is far from complete. In addition to this, complexity increases if an isoperimetric constraint is added to the problem. Such a case is, for example, the problem of extremizing the rise in altitude for a given time. In the present work an attempt is made to treat the problem under the most realistic assumptions used so far, for both the system of equations and the drag model. The analysis of the problem reveals that a more complex thrust history exists than the classical sequence of full-singular-coast subarcs, for both the time-constrained case, and for the case of a drag model with a sharp rise in the transonic region. In the first case, a second full-thrust subarc is generated at the end of the singular subarc, owing to the boundedness of the thrust, while, in the second case, a full-thrust subarc appears in transition from the subsonic to the supersonic branch of the singular path. Both are new results, at least for the bounded-thrust case, and the drag law assumed, insofar as the author knows. Discussion is also provided for the limitations of such a switching structure, and it is shown that the composition of an optimal trajectory is heavily dependent on the assumed drag law.<br>Master of Science
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Phillips, Craig Alan. "Energy management for a multiple-pulse missile." Thesis, Virginia Polytechnic Institute and State University, 1986. http://hdl.handle.net/10919/91058.
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A nonlinear programming technique is applied to the optimization of the thrust and lift control histories for missiles. The first problem considered is that of determining the thrust history which maximizes the range of a continuously-variable (non-pulsed) thrust rocket in horizontal lifting flight. The optimal control solution for this problem is developed. The problem is then approximated by a parameter optimization problem which is solved using a second-order, quasi-Newton method with constraint projection. The two solutions are found to compare well. This result allows confidence in the use of the nonlinear-programming technique to solve optimization problems in flight mechanics for which no analytical optimal-control solutions exist. Such a problem is to determine the thrust and lift histories which maximize the final velocity of a multiple-pulse missile. This problem is solved for both horizontal- and elevation-plane trajectories with and without final time constraints. The method is found to perform well in the solution of these optimization problems and to yield substantial improvements in performance over the nominal trajectories.<br>M.S.
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Krolak, Matthew Joseph. "Optimization of a magnetoplasmadynamic arc thruster." Link to electronic thesis, 2007. http://www.wpi.edu/Pubs/ETD/Available/etd-042607-155701/.
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Gruber, J. R. "A study of erosion due to low-energy sputtering in the discharge chamber of the Kaufman ion thruster." Thesis, University of Oxford, 2002. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.249396.
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Hennig, Christian [Verfasser]. "Improvements in thrust and fuel consumption for future jet engines for unmanned aerial vehicles (UAV) using variable cycle technology / Christian Hennig." München : Verlag Dr. Hut, 2017. http://d-nb.info/1139539094/34.
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Botha, Johannes Rudolf. "Design of an RF ion thruster." Thesis, Stellenbosch : Stellenbosch University, 2014. http://hdl.handle.net/10019.1/86267.
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Thesis (MEng)--Stellenbosch University, 2014.<br>ENGLISH ABSTRACT: Recent years have seen a decline in the rate of space exploration due to the inefficiency
of chemical rockets. Therefore alternative fuel efficient propulsion methods are being
sought to enable cost effective deep space exploration. The high fuel efficiency of electric
thrusters enable a spacecraft to travel further, faster and cheaper than any other propulsion
technology available. Thus electric propulsion has become the propulsion of choice
for scientists and engineers. A typically electric thruster contains some sort of electrode
to ionise the propellant. Although this is feasible for short space missions, it becomes impractical
for more ambitious space missions as electrodes erode over time. The alternative
is to ionise the propellant using electromagnetic fields, which eliminates lifespan issues
associated with electrode based thrusters.
In order to examine methods of improving the lifespan and performance of electric thrusters,
this thesis aimed to study the method of microwave discharge ionisation for an electric
thruster. This includes the design of an RF Ion Thruster with extraction and acceleration
grids to generate thrust. A 600 W 2.45 GHz magnetron (obtained from a conventional
microwave oven), coupled to circular TM010 resonant cavity, was used to ionise neutral
argon gas. The process of electron cyclotron resonance (ECR) was used to ensure the
efficient ionisation of a high density plasma. The thrust was achieved with a three-grid
system biased at high voltages to accelerate positively charged argon ions to high exhaust
velocities.
Results yielded the success of the designed electromagnetic based thruster, measuring
approximatively 1.78 mN of thrust with a specific impulse of Isp = 3786 seconds. The
ECR process produced a high plasma density with a plasma absorption rate of approximately
77% of the total input microwave power. The final results obtained were found
to match the predicted results extremely well and resembled results found in literature.
This demonstrates the efficiency of the RF ion thruster that was designed in this project
and the future use in space exploration activities. However, future research needs to be
undertaken on a controlled feedback system that will ensure optimal operating conditions
for maximum performance. In addition, the method of grid-less acceleration needs to be
studied to achieve maximum thrust and specific impulse.<br>AFRIKAANSE OPSOMMING: In onlangse jare het ’n afname in die tempo van die verkenning van die ruimte dit te danke
aan die ondoeltreffendheid van chemiese vuurpyle. Derhalwe moet alternatiewe brandstof
aandrywing metodes ondersoek word, om koste-effektiewe diep ruimte-eksplorasie
moontlik te maak. Die hoë brandstof-doeltreffendheid van elektriese ontbranders stel ’n
ruimtetuig in staat om verder, vinniger en goedkoper te reis as enige ander aandrywing tegnologie
wat tans beskikbaar is. Dus het elektriese aandrywing metodes die aandrywings
keuse vir wetenskaplikes en ingenieurs geword. ’n Tipies elektriese vuurpyl/aandrywer
bevat ’n vorm van elektrode om die brandstof (argon gas) te ioniseer. Alhoewel hierdie
elektrode proses van ionisasie effektief is vir kort ruimte missies, word dit onprakties vir
meer ambisieuse ruimte missies as gevolg van verweering van elektrodes met verloop van
tyd. ’n Alternatief is om die dryfmiddel/brandstof te ioniseer deur gebruik te maak van
elektromagnetiese velde. Die elekromagnetiese velde sal die lewensduur van die vuurpyl
vermeerder deur die verweering van elektrodes, wat geassosieer word met tipiese elektrieses
vuurpyle, te elimineer.
Hierdie tesis se doelwit is om die metode van mikrogolf ontslag ionisasie vir ’n elektriese
vuurpyl/aandrywer te bestudeer om ten einde die lewensduur en doeltreffendheid van elektriese
vuurpyl/aandrywer te ondersoek. Dit sluit in die ontwerp van ’n radio frekewensie
ioon vuurpyl/aandrywer met ’n ontginning en versnelling matriks/rooster om stukrag te
genereer. ’n 2,45 GHz magnetron (verkry vanaf ’n konvensionele mikrogolfoond), gekoppel
aan ’n TM010 resonante holte, was gebruik om neutrale argon gas te ioniseer. Die
proses van elektron siklotron resonansie (ESR) was gebruik om die doeltreffende ionisasie
van ’n hoë digtheid plasma te verseker. Die aandrywing/stukrag was behaal met ’n
drie-matriks-stelsel, bevoordeel deur hoë spannings om die positief-gelaaide argon ione te
versnel. Resultate opgelewer, het die sukses van die ontwerp van ’n elektromagnetiese gebaseerde
vuurpyl/aandrywer met ’n benaderde meting van ongeveer 1.78 mN van stukrag/aandrywing
met ’n spesifieke impuls van Isp = 3786 sekondes bewys. Die ECR proses het ’n hoë
plasma digtheid geproduseer met ’n plasma opname persentasie van ongeveer 77% van
die totale inset mikrogolf energie. Die finale uitslae wat verkry was, het bevind dat die
voorspelde resultate baie goed inpas met resultate in beskikbare literatuur. Dit dui op die doeltreffendheid van die RF ioon vuurpyl/aandrywer wat ontwerp is in hierdie projek vir
die toekomstige gebruik in ruimte eksplorasie-aktiwiteite. Toekomstige navorsing moet
op ’n beheerde terugvoer sisteem onderneem word, wat optimale werktoestande verseker
vir maksimum prestasie. Daarbenewens moet die metode van matriks-lose versnelling
bestudeer word, om maksimum versnelling/stukrag en spesifieke impuls te verseker.
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Magklaras, Georgios Vasilios. "An insider misuse threat detection and prediction language." Thesis, University of Plymouth, 2012. http://hdl.handle.net/10026.1/1024.
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Numerous studies indicate that amongst the various types of security threats, the problem of insider misuse of IT systems can have serious consequences for the health of computing infrastructures. Although incidents of external origin are also dangerous, the insider IT misuse problem is difficult to address for a number of reasons. A fundamental reason that makes the problem mitigation difficult relates to the level of trust legitimate users possess inside the organization. The trust factor makes it difficult to detect threats originating from the actions and credentials of individual users. An equally important difficulty in the process of mitigating insider IT threats is based on the variability of the problem. The nature of Insider IT misuse varies amongst organizations. Hence, the problem of expressing what constitutes a threat, as well as the process of detecting and predicting it are non trivial tasks that add up to the multi- factorial nature of insider IT misuse. This thesis is concerned with the process of systematizing the specification of insider threats, focusing on their system-level detection and prediction. The design of suitable user audit mechanisms and semantics form a Domain Specific Language to detect and predict insider misuse incidents. As a result, the thesis proposes in detail ways to construct standardized descriptions (signatures) of insider threat incidents, as means of aiding researchers and IT system experts mitigate the problem of insider IT misuse. The produced audit engine (LUARM – Logging User Actions in Relational Mode) and the Insider Threat Prediction and Specification Language (ITPSL) are two utilities that can be added to the IT insider misuse mitigation arsenal. LUARM is a novel audit engine designed specifically to address the needs of monitoring insider actions. These needs cannot be met by traditional open source audit utilities. ITPSL is an XML based markup that can standardize the description of incidents and threats and thus make use of the LUARM audit data. Its novelty lies on the fact that it can be used to detect as well as predict instances of threats, a task that has not been achieved to this date by a domain specific language to address threats. The research project evaluated the produced language using a cyber-misuse experiment approach derived from real world misuse incident data. The results of the experiment showed that the ITPSL and its associated audit engine LUARM provide a good foundation for insider threat specification and prediction. Some language deficiencies relate to the fact that the insider threat specification process requires a good knowledge of the software applications used in a computer system. As the language is easily expandable, future developments to improve the language towards this direction are suggested.
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Thrun, Alexander [Verfasser]. "Investigation of the laminar and turbulent combustion behavior in a one-cylinder internal combustion engine / Alexander Thrun." Kiel : Universitätsbibliothek Kiel, 2016. http://d-nb.info/1117540421/34.
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Ogunjobi, Taiwo A. "Computational Study of Ring-Cusp Magnet Configurations that Provide Maximum Electron Confinement." Wright State University / OhioLINK, 2006. http://rave.ohiolink.edu/etdc/view?acc_num=wright1166226698.
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Penkal, Bryan James. "Steps in the Development of a Full Particle-in-Cell, Monte Carlo Simulation of the Plasma in the Discharge Chamber of an Ion Engine." Wright State University / OhioLINK, 2013. http://rave.ohiolink.edu/etdc/view?acc_num=wright1367586856.
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PAISSONI, CHRISTOPHER ANDREA. "Comprehensive approach to electric propulsion for innovative space platforms." Doctoral thesis, Politecnico di Torino, 2021. http://hdl.handle.net/11583/2945189.
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He, JHIH-JHONG, and 何志忠. "Numerical Thermo-Fluids Analysis of High Thrust Engine''s Hush House." Thesis, 2009. http://ndltd.ncl.edu.tw/handle/76647740259186616780.
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碩士<br>逢甲大學<br>航太與系統工程所<br>97<br>The major objective of this research is to develop a numerical simulation process to study an engine hush house and assure the temperature, velocity and pressure of the internal flow and house wall can meet the hush house regulation. We also adopt three different kinds of noise-reducing pipe installed at the augmenter inlet to understand their influences on the flow fields.
Firstly, we create a hush house model included a F100-PW-100 turbojet engine by using ANSYS Design Modeler software. Meantime, in order to simulate the engine’s jet flow and reflection wave in details, we construct this engine with a true scale convergent-divergent nozzle. Then, we apply ANSYS CFX software to analyze the thermo-fluid performances of the hush house. Finally, we compare the results and conclude some suggestions.
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Shehadeh, Ra'fat. "Study of ejector geometry on thrust augmentation for pulse detonation engine ejector systems." 2007. http://www.etda.libraries.psu.edu/theses/approved/WorldWideIndex/ETD-1961/index.html.
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BETTI, BARBARA. "Flow field and heat transfer analysis of oxygen/methane liquid rocket engine thrust chambers." Doctoral thesis, 2012. http://hdl.handle.net/11573/917591.
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This study is devoted to the characterization and analysis of the flow field and heat transfer in oxygen/methane liquid rocket engines. Attention is focused on the hot gas side of the thrust chamber, where highly energetic flows have to be managed en-
suring the safe operation of the thrust chamber and of the entire engine. Different technological solutions to handle such flows are here investigated by means of CFD numerical simulations. As a compromise between details and computational cost, the attention is focused on capturing the basic phenomena involved which drive the main heat transfer processes, to allow full scale engine analysis as support to the engine
design phase. The simplified approaches are defined, verified and validated against experimental data in different rocket engine conditions, such as film cooled and regeneratively cooled thrust chambers, and expander cycle engine thrust chambers with
heat transfer enhancement devices. Hence, parametric analyses are finally carried out for each configuration. Finally, the simplified approaches are adopted in the thermal analysis of the LM-10 MIRA oxygen/methane expander cycle engine thrust chamber.
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Lung-Da, Wang, and 王隆達. "Assessment for the Effects of Reduced Thrust Taking-off on Airport Environment and Engine Performace." Thesis, 2009. http://ndltd.ncl.edu.tw/handle/57438662384310431924.
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碩士<br>開南大學<br>空運管理學系<br>97<br>Global air transportation is growing up constantly. The environmental protection issue of airborne vehicle is paid attention to continuously by the countries all over the world. There are a lot of methods to reduce the carbon accumulation accompanied with flight energy-conservation at present. However, the reduce thrust take-off was worldwide used by airlines for it neither expanding the cost nor changing the air traffic management.
In order to accurately understand the benefit difference of reduced thrust during take-off. This research noted down the materials in the flights of nationality aviation B747-400, adopting Boeing Method 2. The fuel flow method calculated its CO2 and NOx emission. Compared with ICAO standard, the thrust of reducing has benefits very much on NOx decrement, but in the fuel consumption (CO2 discharged) is not so good as expectancy on quantity. Especially for long haul flight, CO2 emission is greater than ICAO standard.
Secondly, the calculation of the influence on the noise in this research adopts the surveying of the experience formula of SAE AIR 1845 report. The NPD (Noise - power - Distance) form was referred to the INM (Integrated Noise Module) Database. The result shown that reduced thrust can lower noise really, and also discovered that the airplane climbs trajectory (the speed, height change) makes an important factor of degree to decrease the influence. The materials can reflect in the Flight Data Recorder only.
Another important benefit for reduced thrust is to lower the engine hot sections life time and to reduce the maintenance. The research goes through materials by using engine reduced thrust historic data to analysis the trend of engine Exhaust Gas Temperature Margin (EGTM). And, find the thrust of reducing can really make the decline of engine EGTM slow down, and then lengthen the life-span of the engines on wing.
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Zheng, Bo-Sheng, and 鄭博升. "The Develpoment of an Engine Powered Ornithopter- The Measurement of Lift and Thrust of the Flapping-Wing -." Thesis, 2004. http://ndltd.ncl.edu.tw/handle/73026875779671131993.
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碩士<br>大同大學<br>機械工程學系(所)<br>93<br>This research takes designing the engine motive to rush toward the wing machine as the target. By nature the birds fly the observation result of the type, designing the different wing type and the plane wing of the area, and carrying on the measurement that rises the power and pushes the power, discussing in the different wind velocity and offending the Cape under, it rises the power, rushes toward the wing frequency and rises the relation of the power efficiency, finding out the best wing type. The engine motive that then designs one big wing exhibition rushes toward the wing machine, and does the wind tunnel experiment to rise the power and push the power in order to it.
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Zheng, Bo-Sheng, and 鄭博升. "The Development of an Engine Powered Ornithopter ─The Measurement of Lift and Thrust of the Flapping-Wing─." Thesis, 2004. http://ndltd.ncl.edu.tw/handle/84201848984932868160.
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Dores, Delfim Zambujo das Collins Emmanuel G. Alvi Farruk S. "Feedback control for counterflow thrust vectoring with a turbine engine experiment design and robust control design and implementation /." Diss., 2005. http://etd.lib.fsu.edu/theses/available/etd-04082005-130006.
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Thesis (Ph. D.)--Florida State University, 2005.<br>Advisors: Dr. Emmanuel G. Collins Jr., and Dr. Farruk S. Alvi, FAMU-FSU College of Engineering, Dept. of Mechanical Engineering. Title and description from dissertation home page (viewed June 15, 2005). Document formatted into pages; contains xvii, 185 pages. Includes bibliographical references.
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Lee, Chia-Sung, and 李家崧. "Two-Dimensional Thrust-Vectoring Nozzle Design and Dynamic Model Development for a Micro Turbojet Engine and Its Stability Analysis and Verification of a Horizontal Pendulum." Thesis, 2004. http://ndltd.ncl.edu.tw/handle/72632217690701501669.
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碩士<br>中華大學<br>機械與航太工程研究所<br>92<br>A two-dimensional thrust-vectoring nozzle is designed for a P-80 micro turbojet engine. The dynamic model of the side force of this thrust-vectoring engine is validated through the equilibrium of a horizontal pendulum. First, a test stand is developed to measure the axial and the lateral response of the thrust-vectoring nozzle. The performance of the nozzle is used to determine the length of the side wall. The 60mm length of the side wall gives the best steady-state response overall. The turning angle of the side wall is limited to 20 degree in order to have a reasonable exhaust temperature as well as a linear output of the side force. The maximum drop in the axial force is only 6﹪at the maximum turning angle. Then, the dynamic response of this thrust-vectoring engine is examined through a series of different frequency operation. The response of this engine by controlling the fuel flow rate can be modeled as a first order system. The time delay is 0.45 sec and the system bandwidth is 0.175Hz. The response of this engine by turning the angle of the nozzle resembles a first order system except a slight overshoot for frequency between 1.6Hz to 3.18Hz. The time delay is 0.08 sec and the system bandwidth is 2.88Hz by controlling the turning angle. The step response and the ramp response of the axial force and the side force show a good agreement between the models and the test data. The best feedback parameters are 0.05 for the proportional controller, 0.1 for the integral controller, and 1 for the derivative controller for the horizontal pendulum system. The average marginal amplitude is , the steady-state error is , and the oscillating period is 0.95 sec for the turning angle feedback. Finally, several designed parameters are investigated to reduce the marginal amplitude of the horizontal pendulum. Increasing the polar inertia reduces the effect of disturbance. Increasing the damping ratio of the bearing decreases the interference due to viscous effect but the oscillation frequency will increase. For a constant feedback turning angle of the nozzle, there is an optimal distance between the position of the side force and the pendulum support. The correction of the moment is undershoot or overshoot once deviating from the optimum length. Without the disturbance the smaller the feedback turning angle of the nozzle, the smaller the marginal amplitude is. However, the test data shows a minimal turning angle should be used to balance the horizontal pendulum. The interference from the engine itself will make the system unstable for a nozzle correction of less than turning angle. For a nozzle correction of more than turning angle, the larger the marginal amplitude is.
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Терещенко, Юрій Юрійович, Yurii Yurioivych Тereshchenko та Юрий Юрьевич Терещенко. "Концепція інтеграції силової установки з турбовентиляторною приставкою і літального апарату". Thesis, 2019. http://er.nau.edu.ua/handle/NAU/40466.
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У роботі науково обґрунтована концепція аеротермогазодинамічної інте-грації триконтурного газотурбінного двигуна з турбовентиляторною пристав-кою та ступінчастої мотогондоли авіаційної силової установки. Створено нау-ково-методичний апарат для аналізу процесу аеротермогазодинамічної інтегра-ції багатоконтурних турбореактивних двигунів із заднім розташуванням турбо-вентиляторної приставки і мотогондоли авіаційної силової установки. Створено теоретичні основи аеротермогазодинамічної інтеграції багатоконтурного тур-бореактивного двигуна з турбовентиляторою приставкою та мотогондоли сило-вої установки із управлінням примежовим шаром на поверхні мотогондоли га-зогенераторного модуля.
На основі результатів розрахунково-експериментальних досліджень отримані рекомендації щодо обгрунтування вимог до аеротермогазодинамічної інтеграції ступінчастої мотогондоли та газотурбінного двигуна з турбовентиля-торною приставкою авіаційної силової установки та визначення оптимальних значень параметрів робочого процесу триконтурного газотурбінного двигуна відповідно до розрахункових умов польоту.<br>В работе научно обоснована концепция аэротермогазодинамической ин-теграции трехконтурного газотурбинного двигателя с турбовентиляторной при-ставкой и cтупенчатой мотогондолы авиационной силовой установки. Создан научно-методический аппарат для анализа процесса аэротермогазодинамичес-кой интеграции многоконтурных турбореактивных двигателей с задним распо-ложением турбовентиляторной приставки и мотогондолы авиационной силовой установки. Созданы теоретические основы аэротермогазодинамической интег-рации многоконтурного турбореактивного двигателя с турбовентиляторной приставкой и мотогондолы силовой установки с управлением пограничного слоя на поверхности мотогондолы газогенераторного модуля.
Впервые создана методика использования управления пограничным сло-ем на поверхности мотогондолы газогенератора с целью снижения внешнего аэродинамического сопротивления авиационной силовой установки и улучше-ние тягово-экономических характеристик авиационных двухконтурных и трех-контурных турбореактивных двигателей. Впервые получены общие теоретиче-ские зависимости, характеризующие совместную работу модуля мотогондолы с управлением пограничного слоя и принципиально нового типа газотурбинного двигателя с турбовентиляторной приставкой. Разработаны теоретические осно-вы графоаналитического метода определения оптимальных параметров турбо-вентиляторной приставки и мотогондолы газогенератора двухконтурных и тре-хконтурных газотурбинных двигателей, которые позволяют решать задачи оп-тимизации параметров турбовентиляторной приставки.
На основе результатов расчетно-экспериментальных исследований полу-чены рекомендации по обоснованию требований к аэротермогазодинамической интеграции параметров мотогондолы и газотурбинного двигателя с турбовен-тиляторной приставкой авиационной силовой установки и определения опти-мальных значений параметров рабочего процесса трехконтурного газотурбин-ного двигателя в соответствии с расчетными условий полета.<br>The concept of aerothermogasdynamic integration of a double bypass gas tur-bine engine with a turbofan unit and a stepped engine nacelle of an aircraft power plant is scientifically grounded in the work. A scientific and methodological appara-tus has been created for analyzing the process of aerothermogasdynamic integration of multi-pass turbojet engines with a rear-mounted turbofan unit and engine nacelle of an aircraft power plant. The theoretical foundations of aerothermogasdynamic in-tegration of a multi-pass turbojet engine with a turbofan unit and a power plant na-celle with boundary layer control on the surface of the engine nacelle of a gas genera-tor module have been created.
Based on the results of computational and experimental studies, recommenda-tions were obtained to justify the requirements for aerothermogasdynamic integration of a stepped engine nacelle and gas turbine engine with a turbofan unit of an aircraft power plant and determine the optimal values of the process parameters of a of dou-ble bypass gas turbine engine in accordance with the calculated flight conditions.