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1
Chan, Nicholas Y. S. "Scaling considerations for small aircraft engines." Thesis, Massachusetts Institute of Technology, 2008. http://hdl.handle.net/1721.1/45236.
Full textAbstract:
Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2008.Includes bibliographical references (p. 81-84).
Small aircraft engines traditionally have poorer performance compared to larger engines, which until recently, has been a factor that outweighed the aerodynamic benefits of commoditized and distributed propulsion. Improvements in the performance of small engines have, however, prompted another look at this old concept. This thesis examines aspects of aircraft engines that may have application to commodity thrust or distributed propulsion applications. Trends of engine performance with size and time are investigated. These trends are further extended to justify parameter choices for conceptual engines of the current, mid-term (10 years) and far-term (20 years). Uninstalled and installed performances are evaluated for these engines, and parametric studies are performed to determine the most influential and limiting factors. It is found that scaling down of engines is detrimental to SFC and fuel burn, mainly due to the Reynolds number effect. The more scaling done, the more prominent the effect. It is determined that new technology such as higher TIT, OPR and turbomachinery [eta]poly's for small aircraft engines enable the operation of larger bypass ratios, which is the most influential parameter to SFC and fuel bum. The increase of bypass ratio up to a value of 8 is found to be effective for such improvement. SFC decrease from the current to mid-term model is found to be ~20% and ~9% from mid-term to far-term. Range and endurance improvements are found to be ~30% and ~10% respectively for the mission examined. Finally, the mid-term engine model has performance comparable to that of a current, larger state-of-the-art engine, thus suggesting that improvement in small gas turbine technology in the next 10 years will make the application of commodity thrust or distributed propulsion an attractive option for future aircraft.
by Nicholas Y.S. Chan.
S.M.
2
Lee, Victoria D. Lee (Victoria Dawn). "Waste heat reclamation in aircraft engines." Thesis, Massachusetts Institute of Technology, 2014. http://hdl.handle.net/1721.1/97318.
Full textAbstract:
Thesis: S.M., Massachusetts Institute of Technology, Department of Mechanical Engineering, 2014.Cataloged from PDF version of thesis.
Includes bibliographical references (pages 94-96).
Introduction: Rotorcraft engines can lose up to 70% of the potential chemical energy of their fuel as waste heat. Harvesting this waste heat and converting it to useful work would improve the efficiency and power output of the engine. Figure 1 shows two possible engine systems in which a secondary engine could be used to harvest waste heat. For the gas turbine engine in Figure 1A, the main source of waste heat is the enthalpy of the engine's exhaust gases. In the case of the spark ignition engine in Figure 1B, there are three sources of waste heat: the enthalpy available in the exhaust gases, the heat rejected by the coolant loop, and the heat rejected by the oil loop. For each engine system, the heat from waste heat engine is rejected to the ambient air. Possible candidate systems for waste heat recovery include closed cycle systems such as the Rankine and Brayton engines. Rankine engines typical use water as a working fluid. The performance of water-based Rankine engines suffer from low pressures in the working fluid at the temperatures of the ambient and, therefore, require large low pressure expanders and condensers to operate efficiently. Organic working fluids have higher vapor pressures and can be used in Rankine engines instead of water. The higher vapor pressures of these fluids allow the use of smaller expanders. However, organic working fluids are limited to temperatures below 250 C, which is substantially lower than the typical temperatures available in the waste streams. Brayton engines can operate at higher temperatures using inert gases such as helium and argon as working fluids. In either of these engines, the turbomachinery and heat exchangers must remain leak tight as the working fluid is cycled through at high temperatures and high pressures. As a consequence of this requirement, these cycles will not be considered further in this work. Thermoelectric devices, on the other hand, do not require leak tight passages or turbomachinery. These are compacted and are expected to have a higher reliability since they have no moving parts. These advantages have motivated this study on thermoelectrically-based waste heat engine. For a thermoelectrically-based waste heat engine to be feasible, it must be capable of absorbing and rejecting large amounts of heat in part to compensate for the low efficiencies of thermoelectric materials. It must also be light weight and compact to address concerns of power to weight ratios and space constraints in rotorcraft. Therefore, the waste heat engine must be designed to minimize thermal resistance while also minimizing the mass and volume of the heat exchangers.
by Victoria D. Lee.
S.M.
3
Ford, Sean T. "Aerothermodynamic cycle design and optimization method for aircraft engines." Thesis, Georgia Institute of Technology, 2014. http://hdl.handle.net/1853/53006.
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This thesis addresses the need for an optimization method which can simultaneously optimize and balance an aerothermodynamic cycle. The method developed is be able to control cycle design variables at all operating conditions to meet the performance requirements while controlling any additional variables which may be used to optimize the cycle and maintaining all operating limits and engine constraints. The additional variables represent degrees of freedom above what is needed for conservation of mass and energy in the engine system. The motivation for such a method is derived from variable cycle engines, however it is general enough to use with most engine architectures. The method is similar to many optimization algorithms but differs in its implementation to an aircraft engine by combining the cycle balance and optimization using a Newton-Raphson cycle solver to efficiently find cycle designs for a wide range of engine architectures with extra degrees of freedom not needed to balance the cycle. Combination of the optimization with the cycle solver greatly speeds up the design and optimization process. A detailed process description for implementation of the method is provided as well as a proof of concept using several analytical test functions. Finally, the method is demonstrated on a separate flow turbofan model. Limitations and applications of the method are further explored including application to a multi-design point methodology.
4
Ismail, Ibrahim H. "Simulation of aircraft gas turbine engine." Thesis, University of Hertfordshire, 1991. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.303465.
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Sangwian, Sirirat. "Multivariable Sliding Mode Control for Aircraft Engines." Cleveland State University / OhioLINK, 2011. http://rave.ohiolink.edu/etdc/view?acc_num=csu1315587541.
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Ebel, Kathryn C. "Adaptive Sliding Mode Control for Aircraft Engines." Cleveland State University / OhioLINK, 2011. http://rave.ohiolink.edu/etdc/view?acc_num=csu1323882562.
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7
Mahmoud, Saad M. "Effective optimal control of a fighter aircraft engine." Thesis, Loughborough University, 1988. https://dspace.lboro.ac.uk/2134/7287.
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Typical modem fighter aircraft use two-spool, low by-pass ratio, turbojet engines to provide the thrust needed to carry out the combat manoeuvres required by present-day air warfare tactics. The dynamic characteristics of such aircraft engines are complex and non-linear. The need for fast, accurate control of the engine throughout the flight envelope is of paramount importance and this research was concerned with the study of such problems and subsequent design of an optimal linear control which would improve the engine's dynamic response and provide the required correspondence between the output from the engine and the values commanded by a pilot. A detailed mathematical model was derived which, in accuracy and complexity of representation, was a large improvement upon existing analytical models, which assume linear operation over a very small region of the state space, and which was simpler than the large non-analytic representations, which are based on matching operational data. The non-linear model used in this work was based upon information obtained from DYNGEN, a computer program which is used to calculate the steady-state and transient responses of turbojet and turbofan engines. It is a model of fifth order which, it is shown, correctly models the qualitative behaviour of a representative jet engine. A number of operating points were selected to define the boundaries used for the flight envelope. For each point a performance investigation was carried out and a related linear model was established. By posing the problem of engine control as a linear quadratic problem, in which the constraint was the state equation of the linear model, control laws appropriate for each operating point were obtained. A single control was effective with the linear model at every point. The same control laws were then applied to the non-linear mathematical model adjusted for each operating point, and the resulting responses were carefully studied to determine if one single control law could be used with all operating points. Such a law was established. This led, naturally, to the determination of an optimal linear tracking control law, and a further investigation to determine whether there existed an optimal non-linear control law for the non-linear model. In the work presented in this dissertation these points are fully discussed and the reasons for choosing to find an optimal linear control law for the non-linear model by solving the related two-point, boundary value problem using the method of quasilinearisation are presented. A comparison of the effectiveness of the respective optimal control laws, based upon digital simulation, is made before suggestions and recommendations for further work are presented.
8
Alizadeh, Sohail. "Flowfield prediction of NOx and smoke production in aircraft engines." Thesis, Cranfield University, 1993. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.359437.
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9
Helmick, Daniel Martin. "Engine modeling, control, and synchronization for an unmanned aerial vehicle." Thesis, Georgia Institute of Technology, 1998. http://hdl.handle.net/1853/16750.
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10
Harris, P. K. "Erosion in centrifugal compressor impellers." Thesis, Cranfield University, 1996. http://dspace.lib.cranfield.ac.uk/handle/1826/10622.
Full textAbstract:
An experimental and theoretical study of erosion in centrifugal compressor impellers is
presented. An experimental rig using laser anemometry techniques was employed to create
a database of particle restitution ratios for a range of materials. This data was unique in
that the particle rebound was measured in a quiescent condition where the aerodynamic
effects had been minimised, and also parametric factors not previously available were
included. These values were incorporated into the existing Particle Trajectory Code
developed by Cranfield University and Rolls Royce PLC. The code is used to calculate
the trajectories of discrete particles in three dimensional gas turbine geometries, and the
ensuing erosion. It was modified to include the effects of the periodic boundary
conditions, particle fragmentation, splitter blades, and variations in inlet dust concentration
profile. Flowfield calculations were performed on a Rolls Royce GEM-2 and splittered
GEM-60 impeller, which both represent the high pressure stage of the axial + centrifugal
compression system of GEM engines. A procedure developed by Tourlidakis, for the
analysis of steady viscous flow in high speed centrifugal compressors with tip leakage,
was used to generate the flowfields. The GEM-2 impeller flowfield was analysed at 1009c
speed, and validated with calculations and measurements which had been taken for
previous projects. Simulated erosion data under the same conditions was checked using
practical results obtained in a Rolls Royce PLC Helicopter Engine Environmental
Protection Programme, and good agreement was achieved. In order to provide a qualitative,
experimental assessment of erosion, a GEM-60 impeller was coated with four layers of
paint of different colours. Two sizes of quartz particle, each at three different vane
heights, were then seeded into the impeller while it was run cold at (the maximum) 70%
speed. The erosion patterns generated compared well with the results generated by the
Particle Trajectory Code.
11
Chatelier, Adrien. "Toward the study of combustion instabilities in aircraft engines." Thesis, KTH, Aerodynamik, 2014. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-163912.
Full textAbstract:
The prediction of combustion instabilities is crucial for aircraft engines manufacturers to ensure the reliability and the life span of their engines. To study these phenomena, numerical simulations are more and more often performed. They o er advantages: a wide range of parameters is available, they are safe and inexpensive compared to experiments. This work presents a rst step in the general study of combustion instabilities. Firstly, several tools for acoustics and combustion are validated. The tests are performed on simple cases which have reference solutions from analytical resolution or measurements, such as one dimensional ames for combustion and rectangular duct for acoustics. The parameters of these cases are close to the more complex ones. Then, the dynamics of a laminar ame are studied. Finally, a laboratory scale conguration is explored. Kinetics for a methane-air mixture are validated as well as the TFLES model for combustion. The AVSP code for acoustics correctly determines the eigenmodes of a simple conguration. The TFLES model on a pulsed ame has a signicant impact on its dynamics, depending on the thickening factor and the model used. An unstable mode is found for reacting RANS computations of a laboratory scale conguration. An unstable mode is also predicted by AVSP computations with an active ame model. Even if further work is needed to develop these tools, the rst results indicate that they can quickly yield predictions about combustion instabilities.
12
Schonewald, Roger William. "Marketing strategy for commercial aircraft engines : a case study." Thesis, Massachusetts Institute of Technology, 1989. http://hdl.handle.net/1721.1/14478.
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13
Burgess, C. A. R. "The application of aero gas turbine engine monitoring systems to military aircraft." Thesis, Cranfield University, 1990. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.232816.
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14
BENETHUILLERE, Quentin. "Revision Of The Aircraft Engines Preliminary Design Platform Of First Level." Thesis, KTH, Energiteknik, 2014. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-154208.
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In the highly competitive aerospace industry, engine manufacturers must react very quickly and precisely to any demand emerging from aircraft manufacturers if they want to be positioned on the offer. This is especially true when answering to Requests For Information (RFI) based on preliminary design investigations of first level. In order to reduce the time needed to perform these costly operations while improving the performances achieved, Snecma wishes to develop tools for dimensioning the engine and also for assessing key parameters such as mass, emissions, fuel burn, costs, etc. Unfortunately, the set of tools and the process used at the present time for preliminary design investigations of first level are not sufficient to meet the high standards sought-after by the company in terms of time and performances. As a consequence, efforts must be spent on redefining the whole process and the tools it is based on; here is the mission that has been conferred upon me. Multiple exchanges with performances engineers and specialists allowed to draw the current process for preliminary design investigations of first level and raise all the associated concerns. At the same time, a status of the existing tools (called modules in this report), mainly developed under Excel, has been realised in order to identify the range of action for today's investigations. A prototype has been developed under SDK Python with the aim of proving the feasibility of a solution to a difficulty that shows up in the process for each new investigation: the one of generating the workflow on the optimisation software Optimus. A target process has finally been discussed considering all the information collected, and would allow dividing by five the time needed to perform investigations compare to now. The prototype developed lead to interesting results and this solution could thus probably be integrated in the target process as it would allow saving one day of work for an engineer for each study to be carried out. Solutions have been proposed to all the concerns identified in the process and they will have to be discussed with many actors and investigated further in the near future in order to set the target process that will allow meeting the final objective of answering all types of RFIs emitted by aircraft manufacturer in a very short time with a high level of confidence in the results.
15
Sudol, Eugene G. "Evaluation of aircraft turbine redesigns." Monterey, California : Naval Postgraduate School, 1990. http://handle.dtic.mil/100.2/ADA237599.
Full textAbstract:
Thesis (M.S. in Management)--Naval Postgraduate School, June 1990.Thesis Advisor(s): Carrick, Paul M. Second Reader: Doyle, Richard B. "June 1990." Description based on title screen as viewed on October 16, 2009. DTIC Identifier(s): Jet Engines, Engine Components, Cost Analysis, Gas Turbines, Optimizations, Naval Logistics, Aircraft Maintenance, CIP(Component Improvement Program), Benefits, Redesign, Naval Aircraft, Mean Time Between Failure, Data Bases, Theses. Author(s) subject terms: Aircraft Turbine Engine Redesigns Component Improvement Program. Includes bibliographical references (p. 58-60). Also available in print.
16
Hanumanthan, Hariharan. "Severity estimation and shop visit prediction of civil aircraft engines." Thesis, Cranfield University, 2009. http://dspace.lib.cranfield.ac.uk/handle/1826/7634.
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To sustain in the vibrant field of civil aviation, the aircraft and engine
manufacturers are in the pursuit of delivering efficient systems with the best
economics. In umpteen scenarios of growing interest, engine maintenance cost
due to scheduled maintenance is of importance. The current research is
focused on estimation of the maintenance factors, such as severity and shop
visit rate to study the operational scenarios and concurrent technologies.
The severity, defined as relative engine damage is estimated by blending the
aircraft performance, gas turbine performance, gas turbine design and life
estimation methods towards transforming the thrust variation into life estimates,
reflecting the severity on critical Life Limited Part (LLP) of an aircraft engine.
The Shop Visit Rate (SVR) is predicted based on Exhaust Gas Temperature
(EGT) margin consumption due to gas turbine performance degradation.
The severity studies reveal that Hight Pressure Turbine (HPT) blade and disc
are critical, depicting engine severity. Lower thrust engine severity is dominated
by cyclic damage (low cycle fatigue) and large thrust engines by steady state
damage (creep). The operational factors, take-off derate and Outside Air
Temperature (OAT) have more sensitivity on severity of aircraft engines. The
use of climb derate, reduces the damage on large thrust engines considerably,
especially for three shaft engines. Cooling effectiveness and thermal barrier
coating are important technological factors for reducing the severity level.
The SVR prediction on lower and large thrust engines, depict the take-off EGT
as a source for shop visits, governed by operational parameters such as takeoff
derate, OAT, trip length and engine wash. The engine aging curves are
represented as Weibull distribution based on severity and SVR.
Severity estimation and shop visit prediction methodology, demonstrated
through an integrated tool will serve as a decision making element for
comparing competitive engines, operational strategies and engine technologies.
17
Stitt, Alice C. "A physics-based maintenance cost methodology for commercial aircraft engines." Thesis, Cranfield University, 2014. http://dspace.lib.cranfield.ac.uk/handle/1826/13134.
Full textAbstract:
A need has been established in industry and academic publications to link an engine's maintenance costs throughout its operational life to its design as well as its operations and operating conditions. The established correlations between engine operation, design and maintenance costs highlight the value of establishing a satisfactory measure of the relative damage due to different operating conditions (operational severity). The methodology developed in this research enables the exploration of the causal, physics-based relationships underlying the statistical correlations in the public domain and identifies areas for further investigation. This thesis describes a physics-based approach to exploring the interactions, for commercial aircraft, of engine design, operation and through life maintenance costs. Applying the "virtual-workshop" workscoping concept to model engine maintenance throughout the operating life captures the maintenance requirements at each shop visit and the impact of a given shop visit on the timing and requirements for subsequent visits. Comparisons can thus be made between the cost implications of alternative operating regimes, flight profiles and maintenance strategies, taking into account engine design, age, operation and severity. The workscoping model developed operates within a physics-based methodology developed collaboratively within the research group which encompasses engine performance, lifing and operational severity modelling. The tool-set of coupled models used in this research additionally includes the workscoping maintenance cost model developed and implements a simplified 3D turbine blade geometry, new lifing models and an additional lifing mechanism (Thermo-mechanical fatigue (TMF)). Case studies presented model the effects of different outside air temperatures, reduced thrust operations (derate), flight durations and maintenance decisions. The use of operational severity and exhaust gas temperature margin deterioration as physics based cost drivers, while commonly accepted, limit the comparability of the results to other engine-aircraft pairs as the definition of operational severity, its derivation and application vary widely. The use of a single operation severity per mission based on high pressure turbine blade life does not permit the maintenance to vary with the prevalent lifing mechanism type (cyclic/steady state).
18
Weinberg, Maurice 1950. "An optimized product development process for aircraft gas turbine engines." Thesis, Massachusetts Institute of Technology, 2003. http://hdl.handle.net/1721.1/91786.
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Adetifa, Oluwaseun Emmanuel. "Prediction of supersonic fan noise generated by turbofan aircraft engines." Thesis, University of Southampton, 2015. https://eprints.soton.ac.uk/388030/.
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Prediction of Supersonic Fan Noise Generated by Turbofan Aircraft Engines was focussed on improving the capability of predicting supersonic fan noise from modern high-bypass-ratio turbofan aero-engines. The shift from single core jet engines to highbypass-ratio turbofan engines brought about a reduction in the overall aircraft engine noise principally by reducing the jet-broadband noise. However, this new design meant the size of the fan of a high-bypass-ratio turbofan engine, over subsequent years, has increased in diameter. This increase allowed for the speed of the tips of the fan blades to reach and exceed the speed of sound. At high power engine operation conditions, especially at take-off conditions, the noise levels observed from such engines is very high. A major component of this noise is the supersonic fan noise which is also referred to as buzz-saw noise. Shocks are produced at the fan blade tips at this high power engine operation condition. These shocks propagate upstream, against the inflow, following a helical path dictated by the rotation of the fan. The pressure field produced at the tip of the fan is represented as a series of shock waves and expansion waves. As this pressure field advances, it interacts with the incoming flow and acoustic treatment in the intake duct. The shocks in the pressure field are all unique and are of different amplitudes. This is because the fan blades, although manufactured to tight tolerances, are not perfectly alike. Also, the arrangement of these fan blades on the fan hub will also lead to unavoidable differences among the fan blades. These minute differences are reflected in the amplitudes of the shocks, making each shock slightly different from the others. Shocks in the pressure field propagate with respect to the magnitude of their pressure amplitude. Therefore, the shocks travel at different speeds. In the course of propagation, faster shocks catch up with slower ones, and they merge into a single shock, even as the shocks’ amplitudes are attenuated. The difference in speeds and the interactions among the shocks ensure a transfer of energy among the harmonics of the pressure field. This process is nonlinear; the work in this thesis is focussed on modelling the nonlinear propagation of the shocks pressure pattern. These interactions greatly enhance the lower frequency harmonics of the pressure field shifting the dominance from the blade passage frequency and its harmonics. Further upstream, the dominance of the low frequency harmonics is unmistakable. Subsequently the pressure field is radiated from the aircraft intake duct. The resultant radiated pressure field is that which is perceived by an observer in the far-field. The models presented in this thesis capture the main features of this nonlinear propagation and radiation of the pressure field generated at the fan blade tips, and generates predictions for supersonic fan noise levels in the intake duct and in the far-field. A time domain model named SPRID (Sawtooth Propagation in Rigid Intake Ducts) developed is presented. This model predicts the supersonic fan noise levels in ducts without any acoustic treatment, and has been validated against a benchmark frequency domain nonlinear propagation model (FDNS), and also measured data from a modelscale fan rig test provided by Rolls-Royce PLC. The need to incorporate the effect of acoustic liners in the modelling led to the development of a new model which employs the combined time-frequency domain approach. In this model, the nonlinear propagation of the pressure field is simulated in the time domain, while the acoustic liner effects are implemented in the frequency domain. This model also has been validated with measured data. The combined time-frequency domain prediction method was improved to incorporate more complex features of supersonic fan noise propagation. Features such as the change in duct radius along the duct axis and the consequent change in mean flow speeds, and boundary layer effects on the liner absorption have been included in a more advanced model. The advanced nonlinear model is a more representative model of real aircraft intake duct. Also, a theoretical radiation model (GX-Munt) was utilized to predict supersonic fan noise in the far-field. In this thesis, a whole study of supersonic fan noise, starting from source generation at the fan plane up to the radiation to the farfield is presented. The thesis includes an extensive literature review, research on the generation of a source sawtooth for propagation utilizing measured data, and development of equations for nonlinear propagation in axisymmetric intake ducts. Results of the parametric studies using the advanced nonlinear propagation model reliably show all the effects of nonlinear distortion of the shock waves, variation in intake geometry, flow speeds, and variations in the acoustic liner absorption as a consequence of changes in boundary-layer thickness. Comparisons made against measured data, from modelscale fan rig tests conducted by Rolls-Royce PLC, show good and reasonable agreement. The advanced nonlinear propagation model achieves improved prediction capability for supersonic fan noise.
20
Abu, Talib Abd Rahim. "Detailed investigation of the low-temperature analogy of an aircraft engine standard fire-test." Thesis, University of Oxford, 2003. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.289368.
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Skidmore, F. W., and n/a. "The influence of gas turbine combustor fluid mechanics on smoke emissions." Swinburne University of Technology, 1988. http://adt.lib.swin.edu.au./public/adt-VSWT20070420.131227.
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This thesis describes an experimental program covering the development of certain
simple combustion chamber modifications to alleviate smoke emissions from the
Allison T56 turboprop engines operated by the Royal Australian Air Force.
The work includes a literature survey, smoke emission tests on two variants
of the T56 engine, flow visualisation studies of the combustion system in a
water tunnel and combustion rig tests of a standard combustor and four possible
modifications. The rig tests showed that reductions in smoke emissions of
80% were possible by simple modifications that reduced the primary zone
equivalence ratio and improved mixing in that zone.
22
Roy-Aikins, J. E. A. "A study of variable geometry in advanced gas turbines." Thesis, Cranfield University, 1988. http://hdl.handle.net/1826/3907.
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The loss of performance of a gas turbine engine at
off-design is primarily due to the rapid drop of the major
cycle performance parameters with decrease in power and
this may be aggravated by poor component performance. More
and more stringent requirements are being put on the
performance demanded from gas turbines and if future
engines are to exhibit performances superior to those of
present day: engines, then a means must be found of
controlling engine cycle such that the lapse rate of the
major cycle parameters with power is reduced. In certain
applications, it may be desirable to vary engine cycle with
operating conditions in an attempt to re-optimize
performance.
Variable geometry in key engine components offers the
advantage of either improving the internal performance of a
component or re-matching engine cycle to alter the
flow-temperature-pressure relationships. Either method has
the potential to improve engine performance.
Future gas turbines, more so those for aeronautical
applications, will extensively use variable geometry
components and therefore, a tool must exist which is
capable of evaluating the off-design performance of such
engines right from the conceptual stage. With this in mind,
a computer program was developed which can simulate the
steady state performance of arbitrary gas turbines with or
without variable geometry in the gas path components. The
program is a thermodynamic component-matching analysis
program which uses component performance maps to evaluate
the conditions of the gas at the various engine stations.
The program was used to study the performance of a number
of cycles incorporating variable geometry and it was
concluded that variable geometry can significantly improve
the off-design performance of gas turbines.
23
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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Avram, Remus C. "A UNIFIED NONLINEAR ADAPTIVE APPROACH FOR THE FAULT DIAGNOSIS OF AIRCRAFT ENGINES." Wright State University / OhioLINK, 2012. http://rave.ohiolink.edu/etdc/view?acc_num=wright1332784433.
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Ramunno, Michael Angelo. "Control Optimization of Turboshaft Engines for a Turbo-electric Distributed Propulsion Aircraft." The Ohio State University, 2020. http://rave.ohiolink.edu/etdc/view?acc_num=osu1587657623577243.
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Khatami, Iman. "Free-field inlet / outlet noise identification on aircraft engines using microphones array." Thèse, Université de Sherbrooke, 2014. http://hdl.handle.net/11143/7581.
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Abstract : This thesis considers the discrimination of inlet / exhaust noise of aero-engines in free-field static tests using far-field microphone arrays. Various techniques are compared for this problem, including classical beamforming (CB), regularized inverse method (Tikhonov regularization), LI - generalized inverse beamforming (LI-GIB), clean-PSF, clean-SC and two novel methods which are called hybrid method and clean-hybrid. The classical beamforming method is disadvantaged due to its need for a high number of measurement microphones in accordance with the requirements. Similarly, the inverse method is disadvantaged due to their need of having a priori source information. The classical Tikhonov regularization provides improvements in solution stability, however continues to be disadvantaged due to its requirement of imposing a stronger penalty for undetected source positions. Coherent and incoherent sources are resolved by LI-generalized inverse beamforming (L1-GIB). This algorithm can distinguish the multipole sources as well as the monopoles sources. However, source identification by LI-generalized inverse beamforming takes much time and requires a PC with high memory. The hybrid method is a new regularization method which involves the use of an a priori beamforming measurement to define a data-dependent discrete smoothing norm for the regularization of the inverse problem. Compared to the classical beamforming and the inverse modeling, the hybrid (beamforming regularization) approach provides improved source strength maps without substantial added complexity. Although the hybrid method rather solves the disadvantage of the former methods, the application of this method for identification of weaker sources in the presence of the strong sources isn't satisfactory. This can be explained by the large penalization being applied to the weaker source in the hybrid method, which results in underestimation of source strength for this source. To overcome this defect, the clean-SC method and the proposed clean-hybrid method, which is a combination of the hybrid method and the clean-SC, are applied. These methods remove the effect of the strong sources in source power maps to identify the weaker sources. The proposed methods which represent the main contribution of this thesis show promising results and opens new research avenues. Theoretical study of all approaches is performed for various sources and configurations of array. In order to validate the theoretical study, several laboratory experiments are conducted at Universito de Sherbrooke. The proposed methods have further been applied to the measured noise data from a Pratt & Whitney Canada turbo-fan engine and have been observed to provide better spatial resolution and solution robustness with a limited number of measurement microphones compared to the existing methods.Résumé : La présente thèse étudie la discrimination du bruit d'entrée / de sortie des moteurs d'avion dans des tests statiques en champ libre en utilisant des antennes de microphones en champ lointain. Diverses techniques sont comparées pour ce problème, dont la formation de voie classique (CB), la méthode inverse régularisée (régularisation de Tikhonov), la formation de voies généralisée inverse (L1-GIB), Clean-PSF, Clean-SC et deux méthodes proposées qui s'appellent la méthode hybride et la méthode Clean-hybride. La méthode la formation de voie classique est désavantagée en raison de son besoin de nombreux microphones de mesure. De même, la méthode inverse est désavantagée en raison du besoin d'information a priori sur les sources. La régularisation Tikhonov classique fournit des améliorations dans. la stabilité de la solution; cependant elle reste désavantageuse en raison de son exigence d'imposer une pénalité plus forte pour des positions de source non détectées. Des sources cohérentes et incohérentes peuvent être résolues par la formation de voies généralisée inverse (L1-GIB). Cet algorithme peut identifier les sources multi- polaires aussi bien que les sources monopolaires. Cependant, l'identification de source par la formation de voies généralisée inverse prend beaucoup de temps et exige un ordinateur avec une capacité de mémoire élevée. La méthode hybride est une nouvelle méthode de régularisation qui implique l'utilisation d'un traitement par formation de voie a priori pour définir une norme discrète et dépendante des données pour la régularisation du problème inverse. En comparaison avec la formation de voie classique et la méthode inverse, l'approche hybride (régularisation par formation de voie) fournit des cartographies améliorées d'amplitudes de sources sans aucune complexité supplémentaire substantielle. Bien que la méthode hybride lève les limitations des méthodes classiques, l'application de cette méthode pour l'identification de sources de faible puissance en présence de sources de forte puissance n'est pas satisfaisante. On peut expliquer ceci par la plus grande pénalisation appliquée à la source plus faible dans la méthode hybride, qui aboutit à la sous-estimation de l'amplitude de cette source. Pour surmonter ce défaut, la méthode Clean-SC et la méthode Clean-hybrides proposée qui est une combinaison de la méthode hybride et de Clean-SC sont appliquées. Ces méthodes éliminent l'effet des sources fortes dans les cartographies de puissance de sources pour identifier les sources plus faibles. Les méthodes proposées qui représentent la contribution principale de cette thèse conduisent à des résultats fiables et ouvrent des nouvelles voies de recherche. L'étude théorique de toutes les approches est menée pour divers types de sources et de configurations microphoniques. Pour valider l'étude théorique, plusieurs expériences en laboratoire sont réalisées à Université de Sherbrooke. Les méthodes proposées ont été appliquées aux données de bruit mesurées d'une turbo-soufflante Pratt & Whitney Canada pour fournir une meilleure résolution spatiale des sources acoustique et une solution robuste avec un nombre limité des microphones de mesure comparé aux méthodes existantes.
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Pavelec, Sterling Michael. "The development of turbojet aircraft in Germany, Britain, and the United States : a multi-national comparison of aeronautical engineering, 1935-1946 /." The Ohio State University, 2004. http://www.ohiolink.edu/etd/send-pdf.cgi/Pavelec%20Sterling%20Michael.pdf?acc_num=osu1082396007.
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Eveker, Kevin M. "Model Development for active control of stall phenomena in aircraft gas turbine engines." Diss., Georgia Institute of Technology, 1993. http://hdl.handle.net/1853/12363.
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Goh, Shaun Shiao Sing 1980. "Sustainment of commercial aircraft gas turbine engines : an organizational and cognitive engineering approach." Thesis, Massachusetts Institute of Technology, 2003. http://hdl.handle.net/1721.1/82760.
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Martini, Bastien. "Development and assessment of a soot emissions model for aircraft gas turbine engines." Thesis, Massachusetts Institute of Technology, 2008. http://hdl.handle.net/1721.1/45256.
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Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2008.Includes bibliographical references.
Assessing candidate policies designed to address the impact of aviation on the environment requires a simplified method to estimate pollutant emissions for current and future aircraft gas turbine engines under different design and operating assumptions. A method for NOx and CO emissions was developed in a previous research effort. This thesis focuses on the addition of a soot mechanism to the existing model. The goal is to estimate soot emissions of existing gas turbine engines within soot measurement uncertainties, and then to use the method to estimate the performance of potential future engines. Soot is non-volatile primary particulate matter. In gas turbine engines the size rarely exceeds l [mu]m. The soot is composed almost exclusively of black carbon, is an aggregate of nearly spherical carbon primary particles, and exhibits fractal behavior. Results of other studies regarding soot nucleation, growth, oxidation, and coagulation rates are integrated within a network of perfectly-stirred reactors and shown to capture the typical evolution of soot inside a gas turbine combustor, with soot formed in the early parts of the combustor and then oxidized. The soot model shows promising results as its emissions estimates are within the measurement uncertainties. Nevertheless, model uncertainties are high. They are the consequence of the large sensitivity to input variables. Therefore, the validity of the model is limited to cases with available engine data. More engine data are needed to develop and assess the soot model.
by Bastien Martini.
S.M.
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Nygaard, James Robert. "Understanding the behaviour of aircraft bearing steels under rolling contact loading." Thesis, University of Cambridge, 2015. https://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.709284.
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Grasmeyer, Joel M. III. "Multidisciplinary Design Optimization of a Strut-Braced Wing Aircraft." Thesis, Virginia Tech, 1998. http://hdl.handle.net/10919/36729.
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The objective of this study is to use Multidisciplinary Design Optimization (MDO) to investigate the use of truss-braced wing concepts in concert with other advanced technologies to obtain a significant improvement in the performance of transonic transport aircraft. The truss topology introduces several opportunities. A higher aspect ratio and decreased wing thickness can be achieved without an increase in wing weight relative to a cantilever wing. The reduction in thickness allows the wing sweep to be reduced without incurring a transonic wave drag penalty. The reduced wing sweep allows a larger percentage of the wing area to achieve natural laminar flow. Additionally, tip-mounted engines can be used to reduce the induced drag. The MDO approach helps the designer achieve the best technology integration by making optimum trades between competing physical effects in the design space.
To perform this study, a suite of approximate analysis tools was assembled into a complete, conceptual-level MDO code. A typical mission profile of the Boeing 777-200IGW was chosen as the design mission profile. This transport carries 305 passengers in mixed class seating at a cruise Mach number of 0.85 over a range of 7,380 nmi.
Several single-strut configurations were optimized for minimum takeoff gross weight, using eighteen design variables and seven constraints. The best single-strut configuration shows a 15% savings in takeoff gross weight, 29% savings in fuel weight, 28% increase in L/D, and a 41% increase in seat-miles per gallon relative to a comparable cantilever wing configuration.
In addition to the MDO work, we have proposed some innovative, unconventional arch-braced and ellipse-braced concepts. A plastic solid model of one of the novel configurations was created using the I-DEAS solid modeling software and rapid prototyping hardware.Master of Science
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Donaghy, Kevin Robert. "Fire propagation and heat transfer modelling within the BR710 nacelle for certification purposes." Thesis, Queen's University Belfast, 2000. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.322847.
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Schutte, Jeffrey Scott. "Simultaneous multi-design point approach to gas turbine on-design cycle analysis for aircraft engines." Diss., Atlanta, Ga. : Georgia Institute of Technology, 2009. http://hdl.handle.net/1853/28169.
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Thesis (M. S.)--Aerospace Engineering, Georgia Institute of Technology, 2009.Committee Chair: Mavris, Dimitri; Committee Member: Gaeta, Richard; Committee Member: German, Brian; Committee Member: Jones, Scott; Committee Member: Schrage, Daniel; Committee Member: Tai, Jimmy.
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Herbert, Lael S. (Lael Stefan) 1977. "Designing for reliability, maintainability, and sustainability (RM&S) in military jet fighter aircraft engines." Thesis, Massachusetts Institute of Technology, 2002. http://hdl.handle.net/1721.1/82251.
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Dakhel, Pierre Max. "Modeling of particulate matter creation and evolution in aircraft engines, plumes and particle sampling systems." Thesis, Massachusetts Institute of Technology, 2005. http://hdl.handle.net/1721.1/32452.
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Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2005.Includes bibliographical references (p. 105-110).
Environmental and health concerns have recently led to growing efforts to characterize the exhaust gas composition of aircraft engines. Besides major chemical species (N₂, 0₂, C0₂ and H₂0), aircraft engines also emit other species in much lower concentrations but that may also have significant impacts. Particulate Matter (PM) belongs to this category. This thesis presents a model of the microphysical processes leading to the creation of PM and its subsequent interactions with gas phase chemical species in thermodynamic environments typical of aircraft engines and exhaust plumes at ground level. The effects of the turbine and nozzle of an engine on non-volatile PM emissions are addressed first. Results suggest that limited opportunities exist for the modification of the microphysical properties of the non-volatile PM in these environments, leading to the conclusion that the characteristics of the turbine and nozzle of an aircraft engine have little or no influence on aircraft non-volatile emissions. The analysis is then extended downstream to the case of a plume at ground level. Direct comparisons are made to volatile PM measurements obtained from a recent test (APEX). Time-scale arguments are used to suggest that gas to particle conversion at ground level temperatures is a process too slow for volatile particles to exist before the plume reaches the sampling system and thus little if no modification of the PM characteristics should be measured. However, the residence times and temperatures within the sampling system used in APEX are such that significant modification of the PM characteristics within the sampling system is expected.
(cont.) Recommendations to improve the measuring techniques at ground level include lowering the residence time of gas samples inside the sampling system to avoid too large a modification of the flow microphysical characteristics before it reaches the measuring instruments, and careful monitoring of the temperature of the sample throughout the probe and sampling line.
by Pierre Max Dakhel.
S.M.
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Sivapragasam, M. "Numerical and experimental investigations on multiple air jets in counterflow for generating aircraft gas turbine engine inlet flow distortion patterns." Thesis, Coventry University, 2014. http://curve.coventry.ac.uk/open/items/0ad1d0c2-6693-4c6e-9224-5a2237862074/1.
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The performance of an aircraft gas turbine engine is adversely affected by the non-uniform or distorted flow in the inlet duct. Inlet flow distortion lowers the surge margin of the engine‟s compression system with surge occurring at much lower pressure ratios at all engine speeds. The compressor and/or engine are subjected to ground tests in the presence of inlet distortion to evaluate its performance. The simplest method of simulating inlet distortion during these tests is by installing a distortion screen ahead of the engine on the test bed. The uniform inlet flow to the compressor becomes nonuniform with total pressure loss after passing through the distortion screen. Though the distortion screens offer a number of significant advantages, they have some disadvantages. The air jet distortion system can alleviate many of the operational disadvantages encountered with the conventional distortion screens. The system consists of a number of air jets arranged in a circumferential array in a plane and issuing opposite to the primary air flow entering the engine. The jets interact with the primary stream and cause a local total pressure loss due to momentum exchange. The individual mass flow rates from the jets can be varied to obtain a required total pressure pattern ahead of the compressor at the Aerodynamic Interface Plane (AIP). A systematic study of the flow field of confined, turbulent, incompressible, axisymmetric jet issuing into counterflow is covered in this research programme. The jet penetration length and the jet width are reduced compared to unconfined counterflow and a linear relationship between the velocity ratio and the jet length ceases to be valid. The flow field of a circular compressible turbulent jet and then a system of four jets arranged circumferentially and issuing into a confined counterflow was studied experimentally and numerically. For the four jet system the mass flow rates in the four jets were equal in the first part of the study and in the second part they were unequal. The loss in total pressure due to the jet(s) interacting with the counterflow was quantified by a total pressure loss parameter λp0. The total pressure loss increased with increasing mass flow ratio. The total pressure loss distribution was evaluated at several locations behind the jet injector(s). The total pressure non-uniformity quantified by Distortion Index (DI) was found to be highest at a location just downstream of the jet injector and at far downstream locations low values of DI were observed. From the understanding gained with a single jet and four jets in counterflow a methodology was developed to generate a given total pressure distortion pattern at the AIP. The methodology employs computations to obtain the total pressure distortion at the AIP with quasi-one-dimensional inviscid analysis used as a starting point to estimate the mass flow rate in the jets. The inviscid analysis also provides a direction to the iterative procedure to vary the mass flow rate in the jets at the end of each computational step. The methodology is demonstrated to generate a given total pressure distortion pattern using four jets and is further extended to a larger number of jets, twelve and later twenty jets. The total pressure distortion patterns typical of use in aircraft gas turbine engine testing are generated accurately with a smaller number of jets than reported in the literature.
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Sementi, Joshua Paul. "A study of jet exhaust-wing interaction /." Thesis, Connect to this title online; UW restricted, 2005. http://hdl.handle.net/1773/10002.
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Tegeder, Troy. "Development of an efficient solar powered unmanned aerial vehicle with an onboard solar tracker /." Diss., CLICK HERE for online access, 2007. http://contentdm.lib.byu.edu/ETD/image/etd1723.pdf.
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Rau, Karl F. "An analysis of non-integerizing the aircraft engines Cost Effectiveness Analysis Spreadsheet Model (CEAMOD Version 2.0)." Thesis, Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 1993. http://handle.dtic.mil/100.2/ADA276251.
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Thesis (M.S. in Management) Naval Postgraduate School, December 1993.Thesis advisor(s): Alan W. McMasters ; Katsuaki L. Terasawa. "December 1993." Includes bibliographical references. Also available online.
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Reeves, Ross R. "A user's manual for the Cost Effectiveness Analysis Spreadsheet Model for aircraft engines (CEAMOD Version 2.0)." Thesis, Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 1993. http://handle.dtic.mil/100.2/ADA278042.
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Thesis (M.S. in Management) Naval Postgraduate School, December 1993.Thesis advisor(s): Alan W. McMasters ; Katsuaki L. Terasawa. "December 1993." Includes bibliographical references. Also available online.
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Kennedy, Stefan Andrew. "A computational investigation into the effects of lipskin damage on inlet flow distortion in aircraft engines." Thesis, Queen's University Belfast, 2011. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.557636.
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Damage to the lipskin of an engine nacelle can have a significant impact on engine performance and safety, especially in combination with other sources of inlet flow distortion. While the degradation of performance in aircraft propulsion systems as a result of inlet distortion has been well-studied, the impact of lipskln damage upon inlet distortion levels is not well understood. Current practice can result in the grounding of aircraft in order to perform repairs if the damage exceeds the specified tolerance in size and the position on the lip, with these tolerances varying from aircraft to aircraft. If these tolerances are overly conservative, aircraft operating times may be adversely affected. The effect of lipskin damage upon the performance of the inlet of the CF34-3A engine has been investigated using the commercial Computational Fluid Dynamics Package ANSYS CFX in a nacelle model which included the fan. The fan was included in order to capture the interaction between the blades and the upstream distortion. The fan was found to have a stabilising effect upon the flow upstream of the fan, reducing the inlet distortion due to a redistribution of the flow. Other sources of inlet distortion have also been studied, with the effects of high angles of attack and crosswinds having been modelled in isolation, and combined with cases of lipskin damage. A typical take-off configuration was primarily used to assess the effect of the damage due to this being the most critical segment of the flight envelope. The resulting distortion levels have been compared to the engine manufacturers limits in order to assess the severity of the damage.
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Ezik, Oguz. "Calculation of the actual cost of engine maintenance." View thesis, 2003. http://handle.dtic.mil/100.2/ADA412960.
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Thesis (M.S.)--Air Force Institute of Technology, 2003.Title from title screen (viewed July 1, 2004). "March 2003." Vita. "AFIT/GOR/ENS/03-06." "ADA412960"--URL. Includes bibliographical references (p. 87-90). Also issued in paper format.
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Schoch, Eric J. "A simulation of the I3 to D repair process and sparing of the F414-GE-400 jet aircraft engine." Thesis, Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 2003. http://library.nps.navy.mil/uhtbin/hyperion-image/03sep%5FSchoch.pdf.
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Thesis (M.S. in Operations Research)--Naval Postgraduate School, September 2003.Thesis advisor(s): Arnold H. Buss, Kevin J. Maher. Includes bibliographical references (p. 147-148). Also available online.
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Sandlund, Erik. "Modelling the Magnetic Influence of a Jet Aircraft : A study on the magnetic interference of an aircraft configuration and its effect on a magnetometer." Thesis, Linköpings universitet, Teoretisk Fysik, 2019. http://urn.kb.se/resolve?urn=urn:nbn:se:liu:diva-157743.
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Aircraft have been used for the detection of submarines since World War II. The basic concept is to attach a sensor to the back of an aircraft. Since the aircraft is a moving metallic object, it is bound to generate a great deal of interference. Because of this, mathematical models and software have been developed to help filter out this interference and thus make the detection of the submarine easier. Normally, the engines of the aircraft are placed on the wings, quite far away from the sensor. However, for a maritime patrol system in development, the jet engines are placed at the rear of the airframe, generating the necessity to study whether or not they affect the performance of the sensor, which is the purpose of this thesis. Several models were created, tested and simulated for the airframe and jet engines. One of each of these were then combined to create a simulation model for the complete aircraft. A jet engine model that included rotating machinery -- a possible source of magnetic interference -- was also created, but could not be added to the model for the complete aircraft. The magnetic interference was mathematically compensated for, removing the static interference, but not the interference during manoeuvres. The jet engine part of the complete aircraft model did not seem to generate a significant amount of magnetic interference compared to the airframe. An electric dipole, representing a submarine, was then added to the simulation. The data from that simulation was put through the mathematical model and distortions of a few~nT were noticeable during straight courses. The jet engine model that included rotating machinery yielded different results compared to the jet engine model in the complete aircraft model. They seemed to contain signals of higher frequency, which were however not detected by a frequency domain study or present during straight courses. It was thus concluded that using this particular engine model the submarine could probably still be detected if the course of the aircraft was kept straight, though further research is needed with more advanced models for the engine, in particular with regards to the rotating machinery.
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Muller, Yannick. "Coupled thermomechanical fluid-structure interaction in the secondary air system of aircraft engines : contribution to an integrated design method." Valenciennes, 2009. http://ged.univ-valenciennes.fr/nuxeo/site/esupversions/94032a6b-3a17-4aaf-b07a-ce560f117b33.
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Dans un turboréacteur, le système d'air secondaire remplit de multiples fonctions. Les flux d'air secondaire contrôlent les températures des matériaux et l'expansion thermale des parties moteurs, en particulier les écartements des joints d'étanchéité. Pour s'assurer de la réalisation des diverses fonctions dès la phase de développement, les différentes propriétés du gaz doivent être correctement prédîtes. Actuellement, les calculs aérodynamiques, livrant les flux les températures et les pressions d'air, sont séparés des calculs thermiques, livrant les températures matériaux. Les interactions dont le traitement nécessite de nombreuses itérations sont ignorées. En effet, un changement de température matériau modifie l'expansion relative des parties moteurs, redéfinissant ainsi l'écartement des joints qui a son tour contrôle les débits d'air. La définition de l'écartement de joint influant de manière importante sur les pertes de charges, un fort effet de couplage est attendu. Le but de l'étude est de prendre en compte ces interactions au sein d'un nouvel outil combinant analyse du système d'air secondaire et calculs thermique et mécaniques. Une série de modules intégrés permet de considérer ces effets dans les cas stationnaires. Un réseau constitue de nodes représentant les chambres connectées par des éléments assimiles a des pertes de charges constitue la base du concept. Utilisant une formulation compatible avec la topologie Elément Finis, le réseau est imbrique dans le modèle Eléments Finis thermomécanique au sein d'un modèle unique et résolu grâce au logiciel CalculiX. Températures, pressions et débits sont calculés basé sur les températures et déformations matériau de l'itération précédente et servent de conditions limites au calcul thermomécanique dans l'itération suivanteIn jet engines, the secondary air system, or SAS, takes care of a variety of important functions. In particular, secondary air flows control material temperatures and thermal expansion of engine parts, especially seal clearances. To check the fulfilment of these functions in the engine design phase, gas properties, temperatures, pressures and mass flow rates, must be accurately predicted. Up to now, the aerodynamic calculations leading to mass-flow rates, fluid pressures and temperatures and the thermal calculations yielding material temperatures are performed separately. A lot of interactions are neglected, the treatment of which would require numerous time consuming iterations. Indeed, material temperature changes lead to a modification of the expansion of the interacting parts yielding significant modifications in the gaps which control mass-flow rates. Since gap width has an important influence on the pressure losses, the interaction between aerodynamic, thermal and solid mechanics solution to the problem is expected to be important. The present investigation aims at taking this interaction into account in a robust analysis tool, combining SAS, thermal and mechanical analysis. An integrated program suite has been created, which allows to calculate these effects steady state. The basic concept is a network consisting of nodes representing the chambers and connected by pressure loss elements. Using a finite-element-compatible formulation, the network is embedded in a thermo-mechanical finite element model of the engine within an unique model and solved using the free software finite element CalculiX. This is done in the form of a module in which the gas pressure temperature and mass-flow are calculated based on the structural temperature and deformation of the previous iteration and serve as boundary conditions to the thermo-mechanical model for the next iteration
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Outirba, Bilal. "Experimental study of the performance and endurance of carbon fiber brush seals for aero-engines bearing chambers." Doctoral thesis, Universite Libre de Bruxelles, 2017. http://hdl.handle.net/2013/ULB-DIPOT:oai:dipot.ulb.ac.be:2013/258495.
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Over the last decades, it has been progressively acknowledged that reducing the specific fuel consumption and the emission of pollutants as well as improving the thrust-to-weight ratio involves extensive research on advanced sealing technologies. Amongst these, brush seals are particularly well considered for their excellent leakage performance, their low friction properties, and their ability to cope with inevitable rotor excursions during flights. This thesis presents the experimental work that has been carried on in order to characterize carbon brush seals performance in function of the bristle pack geometry and the operating conditions. The analyzed parameters are the bristle free length, the density, and the inter-platedistance. The work, performed by the ULB in collaboration with French engine manufacturer Safran Aircraft Engines, provides first, a description of the test installation that reproduces accurately the severe working conditions encountered in a bearing chamber. A total of eight samples were submitted to extensive testing, and allowed to perform a qualitative analysis of the main performance indicators of a brush seal: the leakage flow, and the seal torque. Complex phenomena acting on the bristle pack were put in evidence under the effect of differential pressure androtation speed, and oil, which fundamentally deteriorate the leakage performance of a brush seal. Subsequently, performance models were developed through empirical correlations, based on the experimental data. They predict the leakage flow and the seal torque as a function of the geometrical parameters and operating conditions. In addition, hysteresis issues were also addressed, and an IR camera helped investigating the heat generation properties of a brush seal.Brush seal samples were submitted to endurance testing, in order to highlight wear mechanisms, and study the performance degradation with the operating time. Oil plays a major part in extending brush seals operating life, despite the leakage performance degradation. Finally, the correlations developed throughout the PhD thesis were used to develop an optimization process in function of the operating conditions of a modern aero-engine. Ultimately, large savings in air consumption were put in evidence when replacing labyrinth seals by brush seals.Doctorat en Sciences de l'ingénieur et technologie
info:eu-repo/semantics/nonPublished
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Ko, Yan-Yee Andy. "The Role of Constraints and Vehicle Concepts in Transport Design: A Comparison of Cantilever and Strut-Braced Wing Airplane Concepts." Thesis, Virginia Tech, 2000. http://hdl.handle.net/10919/32785.
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The purpose of this study is to examine the multidisciplinary design optimization (MDO) of a strut-braced wing (SBW) aircraft compared to similarly designed cantilever wing aircraft. In this study, four different configurations are examined: cantilever wing aircraft, fuselage mounted engine SBW, wing mounted engine SBW, and wingtip mounted engine SBW. The cantilever wing design is used as a baseline for comparison. Two mission profiles were used. The first called for a 7380 nmi range with a 305 passenger load based on a typical Boeing 777 mission. The second profile was supplied by Lockheed Martin Aeronautical Systems (LMAS) and has a 7500 nmi range with a 325 passenger load. Both profiles have a 0.85 cruise Mach number and a 500 nmi reserve range.
Several significant refinements and improvements have been made to the previously developed MDO code for this study. Improvements included using ADIFOR (Automatic Differentiation for FORTRAN) to explicitly compute gradients in the design code. Another major change to the MDO code is the improvement of the optimization architecture to allow for a more robust optimization process.
During the Virginia Tech SBW study, Lockheed Martin Aeronautical Systems (LMAS) was tasked by NASA Langley to evaluate the results of previous SBW studies. During this time, the original weight equations which were obtained from NASA Langleyâ s Flight Optimization System (FLOPS) was replaced by LMAS proprietary equations. A detailed study on the impact of the equations from LMAS on the four designs was done, comparing them to the designs that used the FLOPS equations. Results showed that there was little difference in the designs obtained using the new equations.
An investigation of the effect of the design constraints on the different configurations was performed. It was found that in all the design configurations, the aircraft range proved to be the most crucial constraint in the design. However, results showed that all three SBW designs were less sensitive to constraints than the cantilever wing aircraft.
Finally, a double-deck fuselage concept was considered. A double deck fuselage configuration would result in a greater wing/strut intersection angle which would, in turn, reduce interference drag at that section. Due to the lack of available data on double deck fuselage aircraft, a detailed study of passenger and cargo layout was done. Optimized design showed that there was a small improvement in takeoff gross weight and fuel weight over the single-deck fuselage SBW results when compared with a similarly designed cantilever wing aircraft.Master of Science
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Kline, Sara E. "An Investigation of the Performance of Compliant Finger Seals for use in Gas Turbine Engines using Navier-Stokes and Reynolds Equation Based Numerical Models and Experimental Evaluation." University of Akron / OhioLINK, 2016. http://rave.ohiolink.edu/etdc/view?acc_num=akron1478984223281402.
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Benegas, Jayme Diego. "Evaluation of the Hybrid-Electric Aircraft Project Airbus E-Fan X." Master's thesis, Aircraft Design and Systems Group (AERO), Department of Automotive and Aeronautical Engineering, Hamburg University of Applied Sciences, 2019. http://d-nb.info/1204685894.
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Purpose - This master thesis evaluates the hybrid-electric aircraft project E-Fan X with respect to its economical and environmental performance in comparison to its reference aircraft, the BAe 146-100. The E-Fan X is replacing one of the four jet engines of the reference aircraft by an electric motor and a fan. A turboshaft engine in the cargo compartment drives a generator to power the electric motor. --- Methodology - The evaluation of this project is based on standard aircraft design equations. Economics are based on Direct Operating Costs (DOC), which are calculated with the method of the Association of European Airlines (AEA) from 1989, inflated to 2019 values. Environmental impact is assessed based on local air quality (NOx, Ozone and Particulate Matter), climate impact (CO2, NOx, Aircraft-Induced Cloudiness known as AIC) and noise pollution estimated with fundamental acoustic equations. --- Findings - The battery on board the E-Fan X it is not necessary. In order to improve the proposed design, the battery was eliminated. Nevertheless, due to additional parts required in the new configuration, the aircraft is 902 kg heavier. The turboshaft engine saves only 59 kg of fuel. The additional mass has to be compensated by a payload reduced by 9 passengers. The DOC per seat-mile are up by more than 10% and equivalent CO2 per seat-mile are more than 16% up in the new aircraft. --- Research limitations - Results are limited in accuracy by the underlying standard aircraft design calculations. The results are also limited in accuracy by the lack of knowledge of some data of the project. --- Practical implications - The report contributes arguments to the discussion about electric flight. --- Social implications - Results show that unconditional praise given to the environmental characteristics of this industry project are not justified.