Dissertations / Theses on the topic 'Aircraft gas-turbines'

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.

Thesis (M. S.)--Aerospace Engineering, Georgia Institute of Technology, 2009.
Committee Chair: Zinn, Ben; Committee Member: Fuller, Thomas; Committee Member: Gaeta, Rick; Committee Member: Jagoda, Jeff; Committee Member: Neumeier, Yedidia

With strict environmental regulations, gas turbine emissions have been heavily constrained. This requires operating conditions wherein thermo-acoustic flame instabilities are prevalent. During this process the combustor acoustics and combustion heat release fluctuations are coupled and can cause severe structural damage to engine components, reduced operability, and inefficiency that eventually increase emissions. In order to develop an engine without these problems, there needs to be a better understanding of the physics behind the coupling mechanisms of this instability. Among the several coupling mechanisms, the “velocity coupling” process is the main focus of this thesis. The majority of literature has treated axisymmetric disturbance fields which are typical of longitudinal acoustic forcing and axisymmetric excitation of ring vortices. Two important non-axisymmetric disturbances are: (1) transverse acoustics, in the case of circumferential modes of a multi-nozzle annular combustor and (2) helical flow disturbances, seen in the case of swirling flow hydrodynamic instabilities. With significantly less analytical treatment of this non-axisymmetric problem, a general framework is developed for three-dimensional swirl-stabilized flame response to non-axisymmetric disturbances. The dynamics are tracked using a level-set based G-equation applicable to infinitely thin flame sheets. For specific assumptions in a linear framework, general solution characteristics are obtained. The results are presented separately for axisymmetric and non-axisymmetric mean flames. The unsteady heat release process leads to an unsteady volume generation at the flame front due to the expansion of gases. This unsteady volume generation leads to sound generation by the flame as a distributed monopole source. A sound generation model is developed where ambient pressure fluctuations are generated by this distributed fluctuating heat release source on the flame surface. The flame response framework is used to provide this local heat release source input. This study has been specifically performed for the helical flow disturbance cases to illustrate the effects different modes have on the generated sound. Results show that the effects on global heat release and sound generation are significantly different. Finally, the prediction from the analytical models is compared with experimental data. First, a two-dimensional bluff-body stabilized flame experiment is used to obtain measurements of both the flow and flame position in time. This enables a local flame response comparison since the data are spatially resolved along the flame. Next, a three-dimensional swirl-stabilized lifted flame experiment is considered. The measured flow data is used as input to the G-equation model and the global flame response is predicted. This is then compared with the corresponding value obtained using global CH* chemilumenescence measurements.

Thesis (Ph.D)--Aerospace Engineering, Georgia Institute of Technology, 2008.
Committee Chair: Seitzman, Jerry; Committee Member: Filatyev, Sergei; Committee Member: Jagoda, Jechiel; Committee Member: Lieuwen, Timothy; Committee Member: Shelton, Samuel; Committee Member: Zinn, Ben. Part of the SMARTech Electronic Thesis and Dissertation Collection.

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.

Gas turbine engines are important thermal machines used in industrial and transportation fields. They convert fuel energy into mechanical power or thrust for aerial and maritime vehicles, as well as generate pneumatic and electrical energy that could be used for a large variety of applications. The constant search for fuel burn savings and low pollutant emissions in aviation demands, along with new hardware and material technologies, highly complex engine control systems to optimize fuel consumption throughout the engine operating envelope, and consequently generate more efficient aircraft, in addition to meet the regulatory requirements in terms of safety and performance. These conflicting objectives normally lead to trade-off solutions which are difficult to precisely estimate given the large number of variables involved, including altitude, Mach number, ambient temperature, power and bleed extraction, among others. Therefore, some decisions to characterize the engine controller still reside on experience from previous designs and, as a result, add subjectivity and increase the potential for wrong parameter selection. These control systems significantly contribute to gas turbine performance increase. In this sense, this work proposes the study, design and analysis of multivariable robust controllers for a particular gas turbine engine. Firstly, an algorithmic approach is applied to design an aircraft gas turbine engine controller in a two-degree-of-freedom configuration, obtaining H-infinity robust stabilization. It introduces an optimized loop shape design procedure, with the use of the Genetic Algorithm (GA), to further improve the control system performance, as well as bring the experience applied by controller designers and engineers to an automated process, when setting the parameters to shape the frequency response of the engine control loops. Secondly, a Linear Quadratic Gaussian (LQG) controller, with the Loop Transfer Recovery (LTR) is developed to allow a comparative analysis. The resulting controllers are evaluated by computer simulations under typical operating conditions and compared against each other. Noise immunity is also verified. The complete system is also evaluated against requirements from the aviation industry for commercial aircraft engines. Finally, robustness is evaluated in a similar engine model by generating uncertain state space models based on the boundaries of its nominal model at extreme operating conditions.

In the development of modern rotorcraft vehicles, many unique challenges emerge due to the highly coupled nature of individual rotorcraft design disciplines therefore, the use of an integrated product and process development (IPPD) methodology is necessary to drive the design solution. Through the use of parallel design and analysis, this approach achieves the design synthesis of numerous product and process requirements that is essential in ultimately satisfying the customers demands. Over the past twenty years, Georgia Techs Center for Excellence in Rotorcraft Technology (CERT) has continuously focused on refining this IPPD approach within its rotorcraft design course by using the annual American Helicopter Society (AHS) Student Design Competition as the design requirement catalyst. Despite this extensive experience, however, the documentation of this preliminary rotorcraft design approach has become out of date or insufficient in addressing a modern IPPD methodology. In no design discipline is this need for updated documentation more prevalent than in propulsion system design, specifically in the area of gas turbine technology. From an academic perspective, the vast majority of current propulsion system design resources are focused on fixed-wing applications with very limited reference to the use of turboshaft engines. Additionally, most rotorcraft design resources are centered on aerodynamic considerations and largely overlook propulsion system integration. This research effort is aimed at bridging this information gap by developing a preliminary turboshaft engine design methodology that is applicable to a wide range of potential rotorcraft propulsion system design problems. The preliminary engine design process begins by defining the design space through analysis of the initial performance and mission requirements dictated in a given request for proposal (RFP). Engine cycle selection is then completed using tools such as GasTurb and the NASA Engine Performance Program (NEPP) to conduct thorough parametric and engine performance analysis. Basic engine component design considerations are highlighted to facilitate configuration trade studies and to generate more detailed engine performance and geometric data. Throughout this approach, a comprehensive engine design case study is incorporated based on a two-place, turbine training helicopter known as the Georgia Tech Generic Helicopter (GTGH). This example serves as a consistent propulsion system design reference highlighting the level of integration and detail required for each step of the preliminary turboshaft engine design methodology.

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.

Thesis (M.S.)--West Virginia University, 2000.
Title from document title page. Document formatted into pages; contains x, 157 p. : ill. (some col.) Includes abstract. Includes bibliographical references (p. 86-87).

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.

This thesis presents a computational fluid dynamic (CFD) study of high pressure gas turbine blade design with different realistic inlet temperature and velocity boundary conditions. The effects of blade shaping and inlet conditions can only be fully understood by considering the aerodynamics and heat transfer concurrently; this is in contrast to the sequential method of blade design for aerodynamics followed by cooling. The inlet boundary conditions to the NGV simulations are governed by the existence of discrete fuel injectors in the combustion chamber. An appreciation of NGV shaping design under engine realistic inflow conditions will allow for an identification of the correct three dimensional shaping parameters that should be considered for design optimisation. The Rolls-Royce efficient Navier-Stokes solver, HYDRA, was employed in all computational results for a transonic turbine stage. The single passage unsteady method based on the Fourier Shape Correction is adopted. The solver is validated under both rich burn (hot steak only) and the case with swirl inlet profiles for aerothermal characteristics; good agreement is noted with the validation data. Post processing methods were used in order to obtain time-averaged results and blade visualisations. Subsequently, a surrogate design optimisation methodology using machine learning combined with a Genetic Algorithm is implemented and validated. A study of the effect of NGV compound lean on stage performance is carried out and contrasted for uniform and rich burn inlets, and subsequently for lean burn. Compound lean is shown to produce a tip uploading at the rotor inlet, which is beneficial for rich burn, but detrimental for lean burn. It is also found that for rich burn, fluid driving temperature is more dominant than HTC in determining rotor blade heat transfer, the opposite sense to the uniform inlet. Also, for a lean burn inlet, there is another role reversal, with HTC dominating fluid driving temperature in determining heat transfer. A novel NGV design methodology is proposed that seeks to mitigate the combined effects of inlet hot streak and swirling flow. In essence, the concept two NGVs in a pair are shaped independently of each other, thus allowing the inlet flow non uniformity to be suitably accommodated. Finally, two numerical NGV optimisation studies are undertaken for the combined hot streak and swirl inlet for two clocking positions; vane impinging and passage aligned. Due to the prohibitive cost of unsteady CFD simulations for an optimisation strategy, a suitable objective function at the NGV exit plane is used to minimise rotor tip heat flux. The optimised shape for the passage case resulted in the lowest tip heat flux distribution, however the optimum shape for the impinging case led to the highest gain in stage efficiency. This therefore suggests that NGV lean and clocking position should be a consideration for future optimisation and design of the HP stage.

The persistent thrust for a cleaner, greener environment has prompted air pollution regulations to be enforced with increased stringency by environmental protection bodies all over the world. This has prompted gas turbine manufacturers to move from non-premixed combustion to lean, premixed combustion. These lean premixed combustors operate quite fuel-lean compared to the stochiometric, in order to minimize CO and NOx productions, and are very susceptible to oscillations in any of the upstream flow variables. These oscillations cause the heat release rate of the flame to oscillate, which can engage one or more acoustic modes of the combustor or gas turbine components, and under certain conditions, lead to limit cycle oscillations. This phenomenon, called thermoacoustic instabilities, is characterized by very high pressure oscillations and increased heat fluxes at system walls, and can cause significant problems in the routine operability of these combustors, not to mention the occasional hardware damages that could occur, all of which cumulatively cost several millions of dollars. In a bid towards understanding this flow-flame interaction, this research works studies the heat release response of premixed flames to oscillations in reactant equivalence ratio, reactant velocity and pressure, under conditions where the flame preheat zone is convectively compact to these disturbances, using the G-equation. The heat release response is quantified by means of the flame transfer function and together with combustor acoustics, forms a critical component of the analytical models that can predict combustor dynamics. To this end, low excitation amplitude (linear) and high excitation amplitude (nonlinear) responses of the flame are studied in this work. The linear heat release response of lean, premixed flames are seen to be dominated by responses to velocity and equivalence ratio fluctuations at low frequencies, and to pressure fluctuations at high frequencies which are in the vicinity of typical screech frequencies in gas turbine combustors. The nonlinear response problem is exclusively studied in the case of equivalence ratio coupling. Various nonlinearity mechanisms are identified, amongst which the crossover mechanisms, viz., stoichiometric and flammability crossovers, are seen to be responsible in causing saturation in the overall heat release magnitude of the flame. The response physics remain the same across various preheat temperatures and reactant pressures. Finally, comparisons between the chemiluminescence transfer function obtained experimentally and the heat release transfer functions obtained from the reduced order model (ROM) are performed for lean, CH4/Air swirl-stabilized, axisymmetric V-flames. While the comparison between the phases of the experimental and theoretical transfer functions are encouraging, their magnitudes show disagreement at lower Strouhal number gains show disagreement.

Random vibration of a bladed-disk assembly is studied. A stochastic model for the excitation is developed. A unique feature of this model is the statistical periodicity of the blade forces called cyclostationary. A random process is called wide sense eyeclostationary and its statistics are periodic in time. Factors like the turbulent nature of the flow around the blades, the variability in their geometry, and their nonuniform deterioration contribute to the uncertainty in the excitation. In periodic structures, like the bladed-disk assembly, small variation in the blade excitation may lead to high variability in the response. The model developed includes both random and deterministic excitation. A comparison of the responses due to the random and the deterministic part shows the significance of taking into account the variability in the blade forces. Therefore the assumption that the blade forces are all equal, used by all methods for vibration analysis of bladed disk assemblies, may lead to erroneous estimates of their response, reliability and expected life. It is shown that the response is a cyclostationary process. Therefore the cyclostationary property is preserved from the input to the output. Furthermore the frequency of the second moment of the response is equal to two times the frequency of the excitation.
Master of Science

Thesis (M.S.)--University of Toledo, 2010.
Typescript. "Submitted to the Graduate Faculty as partial fulfillment of the requirements for the Master of Science Degree in Mechanical Engineering." "A thesis entitled"--at head of title. Title from title page of PDF document. Bibliography: p. 90-94.

Two state of the art finite element reduction techniquespreviously validated against the direct finite element method,one based on classical modal analysis and another based oncomponent mode synthesis, are applied for efficient mistunedfree vibration and forced response analysis of several bladeddisk geometries. The methods are first applied to two testcases in order to demonstrate the differences in computationalefficiency as well as to validate the methods againstexperimental data. As previous studies have indicated, nonoticeable differences in accuracy are detected for the currentapplications, while the method based on classical modalanalysis is significantly more efficient. Experimental data(mistuned frequencies and mode shapes) available for one of thetwo test cases are compared with numerical predictions, and agood match is obtained, which adds to the previous validationof the methods (against the direct finite element method).

The influence of blade-to-blade coupling and rotation speedon the sensitivity of bladed disks to mistuning is thenstudied. A transonic fan is considered with part span shroudsand without shrouds, respectively, constituting a high and alow blade-to-blade coupling case. For both cases, computationsare performed at rest as well as at various rotation speeds.Mistuning sensitivity is modelled as the dependence ofamplitude magnification on the standard deviation of bladestiffnesses. The finite element reduction technique based onclassical modal analysis is employed for the structuralanalysis. This reduced order model is solved for sets of randomblade stiffnesses with various standard deviations, i.e. MonteCarlo simulations. In order to reduce the sample size, thestatistical data is fitted to a Weibull (type III) parametermodel. Three different parameter estimation techniques areapplied and compared. The key role of blade-to-blade coupling,as well as the ratio of mistuning to coupling, is demonstratedfor the two cases. It is observed that mistuning sensitivityvaries significantly with rotation speed for both fans due toan associated variation in blade-to-blade coupling strength.Focusing on the effect of one specific engine order on themistuned response of the first bending modes, it is observedthat the mistuning sensitivity behaviour of the fan withoutshrouds is unaffected by rotation at its resonant condition,due to insignificant changes in coupling strength at thisspeed. The fan with shrouds, on the other hand, shows asignificantly different behaviour at rest and resonant speed,due to increased coupling under rotation. Comparing the twocases at resonant rotor speeds, the fan without shrouds is lessor equally sensitive to mistuning than the fan with shrouds inthe entire range of mistuning strengths considered.

This thesis’scientific contribution centres on themistuning sensitivity study, where the effects of shrouds androtation speed are quantified for realistic bladed diskgeometries. However, also the validation of two finite elementreduction techniques against experimental measurementsconstitutes an important contribution.

A full Strouhal number thermo-acoustic model is purposed for the feedback control of self excited noise driven combustors. The inclusion of time delays in the volumetric heat release perturbation models create unique behavioral characteristics which are not properly reproduced within current low Strouhal number thermo acoustic models. New analysis tools using probability density functions are introduced which enable exact expressions for the statistics of a time delayed system. Additionally, preexisting tools from applied mathematics and control theory for spectral analysis of time delay systems are introduced to the combustion community. These new analysis tools can be used to extend sensitivity function analysis used in control theory to explain limits to control effectiveness in self-excited combustors. The control effectiveness of self-excited combustors with actuator constraints are found to be most sensitive to the location of non-minimum phase zeros. Modeling the non-minimum phase zeros correctly require accurate volumetric heat release perturbation models. Designs that removes non-minimum phase zeros are more likely to have poles in the right hand complex plane. As a result, unstable combustors are inherently more responsive to feedback control.

Turbulent combustion processes are inherently unsteady and, thus, a source of acoustic radiation, which occurs due to the unsteady expansion of reacting gases. While prior studies have extensively characterized the total sound power radiated by turbulent flames, their spectral characteristics are not well understood. The objective of this research work is to measure the flow and acoustic properties of an open turbulent premixed jet flame and explain the spectral trends of combustion noise. The flame dynamics were characterized using high speed chemiluminescence images of the flame. A model based on the solution of the wave equation with unsteady heat release as the source was developed and was used to relate the measured chemiluminescence fluctuations to its acoustic emission. Acoustic measurements were performed in an anechoic environment for several burner diameters, flow velocities, turbulence intensities, fuels, and equivalence ratios. The acoustic emissions are shown to be characterized by four parameters: peak frequency (Fpeak), low frequency slope (beta), high frequency slope (alpha) and Overall Sound Pressure Level (OASPL). The peak frequency (Fpeak) is characterized by a Strouhal number based on the mean velocity and a flame length. The transfer function between the acoustic spectrum and the spectrum of heat release fluctuations has an f^2 dependence at low frequencies, while it converged to a constant value at high frequencies. Furthermore, the OASPL was found to be characterized by (Fpeak mfH)^2, which resembles the source term in the wave equation.

Combustor blowout is a very serious concern in modern land-based and aircraft engine combustors. The ability to sense blowout precursors can provide significant payoffs in engine reliability and life. The objective of this work is to characterize the blowout phenomenon and develop a sensing methodology which can detect and assess the proximity of a combustor to blowout by monitoring its acoustic signature, thus providing early warning before the actual blowout of the combustor. The first part of the work examines the blowout phenomenon in a piloted jet burner. As blowout was approached, the flame detached from one side of the burner and showed increased flame tip fluctuations, resulting in an increase in low frequency acoustics. Work was then focused on swirling combustion systems. Close to blowout, localized extinction/re-ignition events were observed, which manifested as bursts in the acoustic signal. These events increased in number and duration as the combustor approached blowout, resulting an increase in low frequency acoustics. A variety of spectral, wavelet and thresholding based approaches were developed to detect precursors to blowout. The third part of the study focused on a bluff body burner. It characterized the underlying flame dynamics near blowout in greater detail and related it to the observed acoustic emissions. Vorticity was found to play a significant role in the flame dynamics. The flame passed through two distinct stages prior to blowout. The first was associated with momentary strain levels that exceed the flames extinction strain rate, leading to flame holes. The second was due to large scale alteration of the fluid dynamics in the bluff body wake, leading to violent flapping of the flame front and even larger straining of the flame. This led to low frequency acoustic oscillations, of the order of von Karman vortex shedding. This manifested as an abrupt increase in combustion noise spectra at 40-100 Hz very close to blowout. Finally, work was also done to improve the robustness of lean blowout detection by developing integration techniques that combined data from acoustic and optical sensors.

Le développement des turbines à gaz d’aviation actuelles et futures est principalement axé sur la sécurité, la performance, la minimisation de la consommation de l’énergie, et de plus en plus sur la réduction des émissions d’espèces polluantes. Ainsi, les phases de design de moteurs sont soumises auxaméliorations continues par des études expérimentales et numériques. La présente thèse se consacre à l’étude numérique des phases transitoires et stationnaires de la combustion au sein d’une turbine à gaz d’aviation opérant à divers modes de combustion. Une attention particulière est accordée à la précision des résultats, aux coûts de calcul, et à la facilité de manipulation de l’outil numérique d’un point de vue industriel. Un code de calcul commercial largement utilisé en industrie est donc choisi comme outil numérique. Une méthodologie de Mécanique des Fluides Numériques (MFN) constituée de modèles avancés de turbulence et de combustion jumelés avec un modèle d’allumage sous-maille, est formulé pour prédire les différentes phases de la séquence d’allumage sous différentes conditions d’allumage par temps froid et de rallumage en altitude, ainsi que les propriétés de la flamme en régime stationnaire. Dans un premier temps, l’attention est focalisée sur le régime de combustion stationnaire. Trois méthodologies MFN sont formulées en exploitant trois modèles de turbulence, notamment, le modèle basé sur les équations moyennées de Navier-Stokes instationnaires (URANS), l’adaptation aux échelles de l’écoulement (SAS), et sur la simulation aux grandes échelles (LES). Pour évaluer la pertinence de l’incorporation d’un modèle de chimie détaillée ainsi que celle des effets de chimie hors-équilibre, deux différentes hypothèses sont considérées : l’hypothèse de chimie-infiniment-rapide à travers le modèle d’équilibre-partiel, et l’hypothèse de chimie-finie via le modèle de flammelettes de diffusion. Pour chacune des deux hypothèses, un carburant à une composante, et un autre à deux composantes sont utilisés comme substituts du kérosène (Jet A-1). Les méthodologies MFN résultantes sont appliquées à une chambre de combustion dont l’écoulement est stabilisé par l’effet swirl afin d’évaluer l’aptitude de chacune d’elle à prédire les propriétés de combustion en régime stationnaire. Par la suite, les rapports entre le coût de calcul et la précision des résultats pour les trois méthodologies MFN formulées sont explicitement comparés. La deuxième étude intermédiaire est dédiée au régime de combustion transitoire, notamment à la séquence d’allumage précédant le régime de combustion stationnaire. Un brûleur de combustibles gazeux, muni d’une bougie d’allumage, et dont la flamme est stabilisée par un accroche-flamme, est utilisé pour calibrer le modèle MFN formulé. Ce brûleur, de géométrie relativement simple, peut aider à la compréhension des caractéristiques d’écoulements réactifs complexes, en l’occurrence l’allumabilité et la stabilité. La méthodologie MFN la plus robuste issue de la précédente étude est reconsidérée. Puisque le brûleur fonctionne en mode partiellement pré-mélangé, le modèle de combustion paramétré par la fraction de mélange et la variable de progrès est adopté avec les hypothèses de chimie-infiniment-rapide et de chimie-finie, respectivement à travers le modèle de Bray-Moss-Libby (BML) et un modèle de flammelettes multidimensionnel (FGM). Le modèle d’allumage sous-maille est préalablement ajusté via l’implémentation des propriétés de la flamme considérée. Par la suite, le modèle d’allumage est couplé au solveur LES, puis successivement aux modèles BML et FGM. Pour évaluer les capacités prédictives des méthodologies résultantes, ces dernières sont utilisées pour prédire les évènements d’allumage résultant d’un dépôt d’énergie par étincelles à diverses positions du brûleur, et les résultats sont qualitativement et quantitativement validés en comparant ceux-ci à leurs homologues expérimentaux. Finalement, la méthodologie MFN validée en configuration gazeuse est étendue à la combustion diphasique en la couplant au module de la phase liquide, et en incorporant les propriétés de la flamme de kérosène dans le modèle d’allumage. La méthodologie MFN résultant de cette adaptation, est préalablement appliquée à la chambre de combustion étudiée antérieurement, pour prédire la séquence d’allumage et améliorer les prédictions antérieures des propriétés de la flamme en régime stationnaire. Par la suite, elle est appliquée à une chambre de combustion plus réaliste pour prédire des évènements d’allumage sous différentes conditions d’allumage par temps froid, et de rallumage en altitude. L’aptitude de la nouvelle méthodologie MFN à prédire les deux types d’allumage considérés est mesurée quantitativement et qualitativement en confrontant les résultats des simulations numériques avec les enveloppes d’allumage expérimentales et les images d’une séquence d’allumage enregistrée avec une caméra infrarouge.
The development of current and future aero gas turbine engines is mainly focused on the safety, the performance, the energy consumption, and increasingly on the reduction of pollutants and noise level. To this end, the engine’s design phases are subjected to improving processes continuously through experimental and numerical investigations. The present thesis is concerned with the simulation of transient and steady combustion regimes in an aircraft gas turbine operating under various combustion modes. Particular attention is paid to the accuracy of the results, the computational cost, and the ease of handling the numerical tool from an industrial standpoint. Thus, a commercial Computational Fluid Dynamics (CFD) code widely used in industry is selected as the numerical tool. A CFD methodology consisting of its advanced turbulence and combustion models, coupled with a subgrid spark-based ignition model, is formulated with the final goal of predicting the whole ignition sequence under cold start and altitude relight conditions, and the main flame trends in the steady combustion regime. At first, attention is focused on the steady combustion regime. Various CFD methodologies are formulated using three turbulence models, namely, the Unsteady Reynolds-Averaged Navier-Stokes (URANS), the Scale-Adaptive Simulation (SAS), and the Large Eddy Simulation (LES) models. To appraise the relevance of incorporating a realistic chemistry model and chemical non-equilibrium effects, two different assumptions are considered, namely, the infinitely-fast chemistry through the partial equilibrium model, and the finite-rate chemistry through the diffusion flamelet model. For each of the two assumptions, both one-component and two-component fuels are considered as surrogates for kerosene (Jet A-1). The resulting CFD models are applied to a swirl-stabilized combustion chamber to assess their ability to retrieve the spray flow and combustion properties in the steady combustion regime. Subsequently, the ratios between the accuracy of the results and the computational cost of the three CFD methodologies are explicitly compared. The second intermediate study is devoted to the ignition sequence preceding the steady combustion regime. A bluff-body stabilized burner based on gaseous fuel, and employing a spark-based igniter, is considered to calibrate the CFD model formulated. This burner of relatively simple geometry can provide greater understanding of complex reactive flow features, especially with regard to ignitability and stability. The most robust of the CFD methodologies formulated in the previous configuration is reconsidered. As this burner involves a partially-premixed combustion mode, a combustion model based on the mixture fraction-progress variable formulation is adopted with the assumptions of infinitely-fast chemistry and finite-rate chemistry through the Bray-Moss-Libby (BML) and Flamelet Generated Manifold (FGM) models, respectively. The ignition model is first customized by implementing the properties of the flame considered. Thereafter, the customized ignition model is coupled to the LES solver and combustion models based on the two above-listed assumptions. To assess the predictive capabilities of the resulting CFD methodologies, the latter are used to predict ignition events resulting from the spark deposition at various locations of the burner, and the results are quantitatively and qualitatively validated by comparing the latter to their experimental counterparts. Finally, the CFD methodology validated in the gaseous configuration is extended to spray combustion by first coupling the latter to the spray module, and by implementing the flame properties of kerosene in the ignition model. The resulting CFD model is first applied to the swirl-stabilized combustor investigated previously, with the aim of predicting the whole ignition sequence and improving the previous predictions of the combustion properties in the resulting steady regime. Subsequently, the CFD methodology is applied to a scaled can combustor with the aim of predicting ignition events under cold start and altitude relight operating conditions. The ability of the CFD methodology to predict ignition events under the two operating conditions is assessed by contrasting the numerical predictions to the corresponding experimental ignition envelopes. A qualitative validation of the ignition sequence is also done by comparing the numerical ignition sequence to the high-speed camera images of the corresponding ignition event.

Gas turbine reheat thrust augmenters known as afterburners are used to provide additional thrust during emergencies, take off, combat, and in supersonic flight of high-performance aircrafts. During the course of reheat development, the most persistent trouble has been the onset of high frequency combustion instability, also known as screech, invariably followed by rapid mechanical failure. The coupling of acoustic pressure upstream of the flame stabilizer with in-phase heat-release downstream, results in combustion instability by which the amplitude at various resonant modes — longitudinal (buzz — low frequency), tangential or radial (screech — high frequency) – amplifies leading to deterioration of the afterburner components. Various researchers in early 1950s have performed extensive testing on straight jet afterburners, to identify screech frequencies. Theoretical and experimental work at test rig level has been reported in the case of buzz to validate the heat release combustion models. In this work, focus is given to study the high frequency tangential combustion instability by vibro-acoustic software and the tests are conducted on the scaled bypass flow afterburner for confirmation of predicted screech frequencies. The wave equation for the afterburner is solved taking the appropriate geometry of the afterburner and taking into account the factors affecting the stability. Nozzle of the afterburner is taken into account by using the nozzle admittance condition derived for a choked nozzle. Screech liner admittance boundary condition is imposed and the effect on acoustic attenuation is studied. A new combustion model has been proposed for obtaining the heat release rate response function to acoustic oscillations. Acoustic wave – flame interactions involve unsteady kinetic, fluid mechanic and acoustic processes over a large range of time scales. Three types of flow disturbances exist such as : vortical, entropy, and acoustic. In a homogeneous, uniform flow, these three disturbance modes propagate independently in the linear approximation. Unsteady heat release also generates entropy and vorticity disturbances. Since flow is not accelerated in the region of uniform area duct, vortical and entropy disturbances are treated as in significant, as these disturbances are convected out into atmosphere like an open-ended tube, but these are considered in deriving the nozzle admittance condition. Heat release fluctuations that arise due to fluctuating pressure and temperature are taken into consideration. The aim is to provide results on how flames respond to pressure disturbances of different amplitudes and characterised by different length scales. The development of the theory is based on large activation energy asymptotics. One-dimensional conservation equations are used for obtaining the response function for the heat release rate assuming the laminar flamelet model to be valid. The estimates are compared with the published data and deviations are discussed. The normalized acoustic pressure variation in the afterburner is predicted using the models discussed earlier to provide an indication of the resonant modes of the pressure oscillations and the amplification and attenuation of oscillations caused by the various processes. Similar frequency spectrum is also obtained experimentally using a test rig for a range of inlet mean pressures and temperatures with combustion and core and bypass flows simulated, for confirmation of predicted results. Without the heat source only longitudinal acoustic modes are found to be excited in the afterburner test section. With heat release, three additional tangential modes are excited. By the use of eight probes in the circumferential cross section of afterburner it was possible to identify the tangential modes by their respective phase shift in the experiments. Comparison of normalized acoustic pressure and phase with and without the incorporation of perforate liner is made to study the effectiveness of the screech liner in attenuating the amplitude of screech modes. By the analysis, conclusion is drawn about modes that get effectively attenuated with the presence of perforate liner. Parametric study of screech liner porosity factor of 1.5 % has not shown appreciable attenuation. Whereas with 2.5 % porosity significant attenuation is noticed, but with 4 % porosity, the gain is very minimal. Hence, the perforate screech liner with the porosity of 2.5 % is finalized. From the rig runs, first pure screech tangential mode and second screech coupled tangential modes are captured. The theoretical frequencies for first and second tangential modes with their phases are comparable with experimental results. Though third tangential mode is predicted, it was not excited in the experiments. There was certain level of deviation in the prediction of these frequencies, when compared to the experimentally obtained values. For this test section of length to diameter ratio of 5, no radial modes are encountered both in the analysis and experiments in the frequency range of interest. In summary, an acoustic model has been developed for the afterburner combustor, taking into account the combustion response, the screech liner and the nozzle to study the acoustic instability of the afterburner. The model has been validated experimentally for screech frequencies using a model test rig and the results have given sufficient confidence to apply the model for full scale afterburners as a predictive design tool.

Gas turbine reheat thrust augmenters known as afterburners are used to provide additional thrust during emergencies, take off, combat, and in supersonic flight of high-performance aircrafts. During the course of reheat development, the most persistent trouble has been the onset of high frequency combustion instability, also known as screech, invariably followed by rapid mechanical failure. The coupling of acoustic pressure upstream of the flame stabilizer with in-phase heat-release downstream, results in combustion instability by which the amplitude at various resonant modes — longitudinal (buzz — low frequency), tangential or radial (screech — high frequency) – amplifies leading to deterioration of the afterburner components. Various researchers in early 1950s have performed extensive testing on straight jet afterburners, to identify screech frequencies. Theoretical and experimental work at test rig level has been reported in the case of buzz to validate the heat release combustion models. In this work, focus is given to study the high frequency tangential combustion instability by vibro-acoustic software and the tests are conducted on the scaled bypass flow afterburner for confirmation of predicted screech frequencies. The wave equation for the afterburner is solved taking the appropriate geometry of the afterburner and taking into account the factors affecting the stability. Nozzle of the afterburner is taken into account by using the nozzle admittance condition derived for a choked nozzle. Screech liner admittance boundary condition is imposed and the effect on acoustic attenuation is studied. A new combustion model has been proposed for obtaining the heat release rate response function to acoustic oscillations. Acoustic wave – flame interactions involve unsteady kinetic, fluid mechanic and acoustic processes over a large range of time scales. Three types of flow disturbances exist such as : vortical, entropy, and acoustic. In a homogeneous, uniform flow, these three disturbance modes propagate independently in the linear approximation. Unsteady heat release also generates entropy and vorticity disturbances. Since flow is not accelerated in the region of uniform area duct, vortical and entropy disturbances are treated as in significant, as these disturbances are convected out into atmosphere like an open-ended tube, but these are considered in deriving the nozzle admittance condition. Heat release fluctuations that arise due to fluctuating pressure and temperature are taken into consideration. The aim is to provide results on how flames respond to pressure disturbances of different amplitudes and characterised by different length scales. The development of the theory is based on large activation energy asymptotics. One-dimensional conservation equations are used for obtaining the response function for the heat release rate assuming the laminar flamelet model to be valid. The estimates are compared with the published data and deviations are discussed. The normalized acoustic pressure variation in the afterburner is predicted using the models discussed earlier to provide an indication of the resonant modes of the pressure oscillations and the amplification and attenuation of oscillations caused by the various processes. Similar frequency spectrum is also obtained experimentally using a test rig for a range of inlet mean pressures and temperatures with combustion and core and bypass flows simulated, for confirmation of predicted results. Without the heat source only longitudinal acoustic modes are found to be excited in the afterburner test section. With heat release, three additional tangential modes are excited. By the use of eight probes in the circumferential cross section of afterburner it was possible to identify the tangential modes by their respective phase shift in the experiments. Comparison of normalized acoustic pressure and phase with and without the incorporation of perforate liner is made to study the effectiveness of the screech liner in attenuating the amplitude of screech modes. By the analysis, conclusion is drawn about modes that get effectively attenuated with the presence of perforate liner. Parametric study of screech liner porosity factor of 1.5 % has not shown appreciable attenuation. Whereas with 2.5 % porosity significant attenuation is noticed, but with 4 % porosity, the gain is very minimal. Hence, the perforate screech liner with the porosity of 2.5 % is finalized. From the rig runs, first pure screech tangential mode and second screech coupled tangential modes are captured. The theoretical frequencies for first and second tangential modes with their phases are comparable with experimental results. Though third tangential mode is predicted, it was not excited in the experiments. There was certain level of deviation in the prediction of these frequencies, when compared to the experimentally obtained values. For this test section of length to diameter ratio of 5, no radial modes are encountered both in the analysis and experiments in the frequency range of interest. In summary, an acoustic model has been developed for the afterburner combustor, taking into account the combustion response, the screech liner and the nozzle to study the acoustic instability of the afterburner. The model has been validated experimentally for screech frequencies using a model test rig and the results have given sufficient confidence to apply the model for full scale afterburners as a predictive design tool.

Aero engines are one of the most complex machines on this planet and have propelled the necessity of advanced material technologies. Health monitoring of the engine is performed by a variety of sensors and amongst these strain sensor is very important as it aids in evaluating the stresses experienced by the body. Unlike conventional foil gauge which tends to debond under hostile environments in the engine like high rpm of blades, temperature, mass flow, etc, thin film based strain gauges are likely to exhibit better adhesion on the substrates. The usage of Ni-Cr thin films in strain gauge sensor has been proven for static application, wherein the substrate does not experience the fluctuating loads. Material and mechanical aspects of thin films for design and development of thin film based strain gauge sensor for aero engine application was taken up as a research work. One of the objectives of the work was to characterize the Ni-Cr thin films with varying composition deposited by sputter deposition process and characterize the films for its microstructural features and mechanical properties. The correlation of these properties is performed and amongst the film compositions investigated the film with alloy composition of Ni-Cr:80-20 at% exhibits the most distinct columnar structure, highest electrical resistivity (2.037 μΩm), hardness (5.8 GPa) and the modulus (180 GPa). This Ni-Cr: 80-20 at% film exhibits no surface cracks when loaded in the elastic region of the titanium alloy GTM-Ti-64. Resistance to deformation under the action of externally applied load on a body results in stress within the body. In single or multilayer film stacking the stress experienced in the film by virtue of substrate deformation needs to be investigated quantitatively. The substrate stresses are transferred to the films by shear stresses at the interface. In order to measure the surface strain by change in the electrical resistance of the gage it is important to quantitatively evaluate the stresses in the films. Are these stresses very high to cause delamination and film cracking or are these stresses too less to be measured. In order to understand the stress evolution and transfer mechanism, an analytical approach, numerical simulation and experimental validation were performed. Thin film strain gauge device architecture has been engineered such that an insulating layer of alumina is deposited on substrate and a sensing layer is deposited on the insulating layer to avoid thermal mismatch and maintain the strain compatibility. A alumina of 45 micron thick alumina layer was successfully deposited on Titanium alloy (GTM-Ti-64) by sputter deposition without any edge delamination and microcracks. Finite Element Analysis (FEA) results showed that the axial and shear stress profiles at the Ti alloy-alumina interfaces for both single and multilayer architecture are similar and higher when compared with the stresses in alumina-NiCr. The shear stress profile for single layer and multilayer architecture follows the modified shear lag model with peak shear stresses at the extremes and peak axial stress at the centre of the film. The axial stresses in the alumina film is found to be significant in both FEA and validated by experimental findings with film fracture strength of 814 MPa. Similarly, the shear stresses were found to be minimal by FEA studies and the experimental finding suggests the film fracture under tensile mode. Complete strain transfer was observed from substrate to these thin films under both tensile and vibratory fatigue, suggesting proper adhesion of the alumina film on the Ti alloy substrate. The maximum strain compatibility of thin film alumina on Ti alloy substrate was found to be 0.22 %. A Goodman correction for the fatigue data under axial mode was performed and on combining the entire fatigue data for R = -1 linear fit was observed across all the data points wherein the Basquin equation was considered for data analysis and the fatigue strength coefficient and exponent are found to be 872.56 and -0.054 for alumina thin film on Ti alloy substrate. Thin film strain gauges (TFSG) with these characteristics were deposited on the compressor rotor blade of one of the typical aero engines. Thick contact pads and a new bonding technique are used for taking the lead wires. The entire multilayer structure with wire bonding was tested under static and dynamic (vibratory fatigue) conditions and TFSG exhibited a reproducible strain when compared against foil based strain gauge under both tension and compression. TFSG device was tested for a duration of 2200 seconds with a blade vibration frequency of 406 Hz i.e. 8.9x105 cycles. During the entire test duration, TFSG successfully measured strains from the aero engine blade.