During gas turbine development testing, measurements of the gas-path stagnation temperature are used to characterise the engine running condition and establish individual engine component performance. These measurements are typically acquired using passively ventilated thermocouple probes, which are capable of achieving absolute stagnation temperature uncertainties of approximately 0.5 %. Historically, this measurement accuracy has been considered adequate to evaluate gains in turbomachinery efficiency. However, realisable turbomachinery efficiency gains have recently become sufficiently small that an improvement in measurement accuracy is now required. This has resulted in the specification of a target absolute stagnation temperature uncertainty of 0.1 %. The research presented in this thesis focusses on the development of a new stagnation temperature probe that will achieve a measurement uncertainty close to the target value. The new probe has been designed to utilise a thin-film platinum resistance thermometer (PRT) as the temperature sensitive element. For certain aspects of gas turbine engine testing, this type of sensor offers an improvement in measurement accuracy compared to a thermocouple.

Thesis -- University of Adelaide, 1966. / [Typescript].

A three-dimensional CFD program was used to predict the flow field that would enter a downstream fan or compressor rotor under the influence of an upstream thermal distortion. Two distortion generation techniques were implemented in the model; (1) a thermal source and (2) a heated flow injection method. Results from the investigation indicate that both total pressure and velocity boundary conditions at the compressor face are made non-uniform by the upstream thermal distortion, while static pressure remains nearly constant. Total pressure at the compressor face was found to vary on the order of 10%, while velocity varies from 50-65%. Therefore, in modeling such flows, neither of these latter two boundary conditions can be assumed constant under these conditions. The computational model results for the two distortion generation techniques were compared to one another and evaluations of the physical practicality of the thermal distortion generation methods are presented. Both thermal distortion methods create total temperature distortion magnitudes at the compressor face that may affect rotor blade vibration. Both analyses show that holding static pressure constant is an appropriate boundary condition for flow modeling at the compressor inlet. The analyses indicate that in addition to the introduction of a thermal distortion, there is a potential to generate distortion in total pressure, Mach number, and velocity. Depending on the method of thermally distorting the inlet flow, the flow entering the compressor face may be significantly non-uniform. The compressor face boundary condition results are compared to the assumptions of a previous analysis (Kenyon et al., 2004) in which a 25 R total temperature distortion was applied to a computational fluid dynamics (CFD) model of a fan geometry to obtain unsteady blade pressure loading. Results from the present CFD analyses predict similar total temperature distortion magnitudes corresponding to the total temperature variation used in the Kenyon analyses. However, the results indicate that the total pressure and circumferential velocity boundary conditions assumed uniform in the Kenyon analyses could vary by the order of 2% in total pressure and approximately 8% in velocity distortion. This supports the previously stated finding that assuming a uniform total pressure profile at the compressor inlet may be an appropriate approximation with the presence of a weak thermal distortion, while assuming a constant circumferential velocity boundary condition is likely not sufficiently accurate for any thermal distortion. In this work, the referenced Kenyon investigation and others related to the investigation of distortion-induced aeromechanical effects in this compressor rotor have assumed no aerodynamic coupling between the duct flow and the rotor. A full computational model incorporating the interaction between the duct flow and the fan rotor would serve to alleviate the need for assuming boundary conditions at the compressor inlet. / Master of Science

The ability to predict accurately the critical quality characteristics of aircraft engines is essential for modeling the degradation of engine performance over time. The acceptable margins for error grow smaller with each new generation of engines. This paper focuses on turbine gas temperature (TGT). The goal is to improve the first principles predictions through the incorporation of the pure thermodynamics, as well as available information from the engine health monitoring (EHM) data and appropriate maintenance records. The first step in the approach is to develop the proper thermodynamics model to explain and to predict the observed TGTs. The resulting residuals provide the fundamental information on degradation. The current engineering models are ad hoc adaptations of the underlying thermodynamics not properly tuned by actual data. Interestingly, pure thermodynamics model uses only two variables: atmospheric temperature and a critical pressure ratio. The resulting predictions of TGT are at least similar, and sometimes superior to these ad hoc models. The next steps recognize that there are multiple sources of variability, some nested within others. Examples include version to version of the engine, engine to engine within version, route to route across versions and engines, maintenance to maintenance cycles within engine, and flight segment to flight segment within maintenance cycle. The EHM data provide an opportunity to explain the various sources of variability through appropriate regression models. Different EHM variables explain different contributions to the variability in the residuals, which provides fundamental insights as to the causes of the degradation over time. The resulting combination of the pure thermodynamics model with proper modeling based on the EHM data yield significantly better predictions of the observed TGT, allowing analysts to see the impact of the causes of the degradation much more clearly. / Doctor of Philosophy / AEM is major civilian aircraft gas turbine engine manufacturer, serving different airliners and airlines. However, one of its newest models has had performance issues; the engines degraded faster than their in-house model had anticipated, leading to more frequent maintenance and causing significant financial losses to the company. The key objectives of our research project are to produce a model that has higher predictive capabilities than AEM's in-house predictive model (DTGT), and develop a model selection algorithm that allows for direct comparisons among models of vastly different architecture. There are three major components to our research: 1) interdisciplinary studies merging the theory of thermodynamics and regression, 2) the sequential modeling, and 3) the modified Mallows's Cp. We propose a layered sequential approach to the regression modeling, where one regression model is followed by another regression on the residuals of the previous model. We also propose the modified Mallows's Cp, a modification of the Mallows's Cp, as a viable model selection criterion. Our results demonstrated that the sequential approach both outperformed the AEM's in-house model and was found to be more useful than the traditional multiple linear regression. Our results also demonstrated that the modified Mallows's Cp prefer smaller number of parameters than other standard model selection criterion without sacrificing predictive capabilities of its models.

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

Radial gas turbines engines have established prominence in the field of small turbomachinery because of their simplicity, relatively high performance and installation features. Thus they have been used in a variety of applications such as generator sets, small auxiliary power units (APu), air conditioning of aircraft cabins and hybrid electric vehicles turbines. The current research describes the design, manufacturing, construction and testing a radial type small gas turbine. The aim was to design and build the engine to drive directly a high-speed permanent magnet alternator running at 60000 rpmand developing a maximum of 60 W. This direct coupling arrangement produces a portable, light, compact, reliable and environment friendly power generator. These features make the generator set very attractive to use in many applications including emergency power generation for hospitals, in areas of natural disasters such as floods and earthquakes, in remote areas that cannot be served from the national grid, oil rigs, and in confined places of limited spaces. It is important to recognize that the design of the main components, that is, the inward flow radial UFR turbines, the centrifugal compressor and the combustion chamber involve consideration of aero-dynamics, thermodynamics, fluid mechanics, stress analysis, vibration analysis, selection of bearings, selection of suitable materials and the requirements for manufacturing. These considerations are all inter-linked and a procedure has been followed to reach an optimum design. This research was divided into three phases: phase I dealt with the complete design of the inward radial turbine, the centrifugal compressor, the power transmission shaft, the selection of combustion chamber and the bearing housing including the selection of bearings. Phase 2 dealt with mechanical consideration of the rotating components that is stress, thermal and vibration analyses of the turbine rotor, the impeller and the rotating shaft, respectively. Also it dealt with the selection of a suitable fuel and oil lubrication systems and a suitable starting system. Phase 3 dealt with the manufacturing of the gas turbine components, balancing the rotating components, assembling the engine and finally commissioning and then testing the engine. The current work in this thesis has put the light on a new design methodology on determining the optimum principal dimensions of the rotor and the impeller. This method, also, has defined the optimum number of blades and the axial length of the rotor and the impeller. Mathematical models linking the performance parameters and the design variables for the turbine and the compressor have been developed to assist in carrying out parametric studies to study the influence of the design parameters on the performance and on each other. Also, a new graphical matching procedure has been developed for the gas turbine components. This technique can serve as a valuable tool to determine the operating range and the engine running line. Furthermore, it would decide whether the gas turbine engine operates in a region of satisfactory compressor and turbine efficiencies.

Les réglementations toujours plus drastiques sur les émissions de polluants ont conduit au développement de systèmes de combustion opérant en régimes pauvres qui sont malheureusement sujet aux instabilités thermo acoustiques. La capacité de la Simulation aux Grandes Echelles (SGE) à simuler des turbines à gaz industrielles complexes de grande puissance est mise en évidence au cours de ce travail de thèse. Tout d’abord, la SGE est appliquée à un brûleur académique et validée par comparaison à des mesures effectuées à l’Université de Berlin ainsi qu’à des simulations SGE effectuées avec OpenFOAM chez Siemens. Afin de déterminer la stabilité de ce bruleur le couplage entre l’acoustique et la combustion est modélisé par l’approche de type fonction de transfert de flamme (FTF). Suite à ces calcules et l’évaluation de la FTF les fluctuations du nombre de swirl sont identifiées comme un paramètre à même de modifier cette réponse de flamme. Après cette première étape de validation, une turbine à gaz industrielle est simulée en SGE pour deux géométries différentes du brûleur et pour deux points de fonctionnement. La FTF issue de ces calculs est peu influencée par les deux points de fonctionnement. A l’inverse, des légères modifications de la géométrie du swirler modifient les caractéristiques de la FTF montrant que plusieurs mécanismes sont en jeu. Ces mécanismes sont identifiés comme étant la vitesse d’entrée, les fluctuations de swirl et les fluctuations de fraction de mélange. Cette dernière est causée par: 1) la pulsation du débit de carburant injecté et 2) la trajectoire fluctuante des jets de carburant. Bien que le swirler soit conçu pour fournir un mélange le plus homogène possible, d’importantes hétérogénéités de mélange à l’entrée de la chambre de combustion sont présentes. Les perturbations de mélange se combinent avec les fluctuations de vitesse (et donc avec les fluctuations de swirl) aboutissant à des résultats de FTF différents. Un modèle étendu pour la FTF reliant le dégagement de chaleur à la vitesse d’entrée et à la fluctuation de fraction de mélange (modèle MISO) se révèle être une bonne solution pour ces systèmes complexes. Une analyse non linéaire montre en outre que l’amplitude de forçage conduit non seulement à une saturation de la flamme, mais aussi à un changement de la réponse de flamme. La saturation de la flamme n’est vérifiée que pour la FTF globale et le gain augmente localement avec une amplitude croissante. Pour ce système on notera enfin que la flamme linéaire, comme la flamme non linéaire, ne sont pas compactes: certaines zones pourtant situées l’une à coté de l’autre, ont des différences significatives de délai de FTF, montrant que certaines parties de la flamme amortissent l’excitation alors que d’autres l’amplifient. / Modern pollutant regulation have led to a trend towards lean combustion systems which are prone to thermo-acoustic instabilities. The ability of Large Eddy Simulation (LES) to handle complex industrial heavy-duty gas turbines is evidenced during this thesis work. First, LES is applied to an academic single burner in order to validate the modeling against measurements performed at TU Berlin and against OpenFoam LES simulations done at Siemens. The coupling between acoustic and combustion is modeled with the Flame Transfer Function (FTF) approach and swirl number fluctuations are identified changing the FTF amplitude response of the flame. Then, an industrial gas turbine is analyzed for two different burner geometries and operating conditions. The FTF is only slightly influenced for the two operating points but slight modifications of the swirler geometry do modify the characteristics of the FTF showing that a simple model taking only into account the flight time is not appropriate and additional mechanisms are at play. Those mechanisms are identified being the inlet velocity, the swirl and the inlet mixture fraction fluctuations. The latter is caused by two mechanisms: 1) the pulsating injected fuel flow rate and 2) the fluctuating trajectory of the fuel jets. Although the diagonal swirler is designed to provide good mixing, effects of mixing heterogeneities at the combustion chamber inlet occur. Mixture perturbations phase with velocity (and hence with swirl) fluctuations and combine with them to lead to different FTF results. Another FTF approach linking heat release to inlet velocity and mixture fraction fluctuation (MISO model) shows further to be a good solution for complex systems. A nonlinear analysis shows that the forcing amplitude not only leads to a saturation of the flame but also to changes of the delay response. Flame saturation is only true for the global FTF and the gain increases locally with increasing forcing amplitude. Both, the linear and the nonlinear flames, are not compact: flame regions located right next to each other exhibited significant differences in delay meaning that at the same instant certain parts of the flame damp the excitation while others feed it.

Finding limit cycles and their stability is one of the central problems of nonlinear thermoacoustics. However, a limit cycle is not the only type of self-excited oscillation in a nonlinear system. Nonlinear systems can have quasi-periodic and chaotic oscillations. This thesis examines the different types of oscillation in a numerical model of a ducted premixed flame, the bifurcations that lead to these oscillations and the influence of external forcing on these oscillations. Criteria for the existence and stability of limit cycles in single mode thermoacoustic systems are derived analytically. These criteria, along with the flame describing function, are used to find the types of bifurcation and minimum triggering amplitudes. The choice of model for the velocity perturbation field around the flame is shown to have a strong influence on the types of bifurcation in the system. Therefore, a reduced order model of the velocity perturbation field in a forced laminar premixed flame is obtained from Direct Numerical Simulation. It is shown that the model currently used in the literature precludes subcritical bifurcations and multi-stability. The self-excited thermoacoustic system is simulated in the time domain with many modes in the acoustics and analysed using methods from nonlinear dynamical systems theory. The transitions to the periodic, quasiperiodic and chaotic oscillations are via sub/supercritical Hopf, Neimark-Sacker and period-doubling bifurcations. Routes to chaos are established in this system. It is shown that the single mode system, which gives the same results as a describing function approach, fails to capture the period-$2$, period-$k$, quasi-periodic and chaotic oscillations or the bifurcations and multi-stability seen in the multi-modal case, and underpredicts the amplitude. Instantaneous flame images reveal that the wrinkles on the flame surface and pinch off of flame pockets are regular for periodic oscillations, while they are irregular and have multiple time and length scales for quasi-periodic and chaotic oscillations. Cusp formation, their destruction by flame propagation normal to itself, and pinch-off and rapid burning of pockets of reactants are shown to be responsible for generating a heat release rate that is a highly nonlinear function of the velocity perturbations. It is also shown that for a given acoustic model of the duct, many discretization modes are required to capture the rich dynamics and nonlinear feedback between heat release and acoustics seen in experiments. The influence of external harmonic forcing on self-excited periodic, quasi-periodic and chaotic oscillations are examined. The transition to lock-in, the forcing amplitude required for lock-in and the system response at lock-in are characterized. At certain frequencies, even low-amplitude forcing is sufficient to suppress period-$1$ oscillations to amplitudes that are 90$\%$ lower than that of the unforced state. Therefore, open-loop forcing can be an effective strategy for the suppression of thermoacoustic oscillations. This thesis shows that a ducted premixed flame behaves similarly to low-dimensional chaotic systems and that methods from nonlinear dynamical systems theory are superior to the describing function approach in the frequency domain and time domain analysis currently used in nonlinear thermoacoustics.

Multiple high-fidelity, time-accurate computational fluid dynamics simulations were performed to investigate the effects of upstream stator loading and rotor shock strength on vortex shedding characteristics in a single stage transonic compressor. Various configurations of a transonic axial compressor stage, including three stator/rotor axial spacings of close, mid, and far in conjunction with three stator loadings of decreased, nominal, and increased were simulated in order to understand the flow physics of transonic blade-row interactions. Low-speed compressors typically have reduced stator/rotor axial spacing in order to decrease engine weight, and also because there is an increase in efficiency with reduced axial spacing. The presence of a rotor bow shock in high-speed compressors causes additional losses as the shock interacts with the upstream stator trailing edge. This research analyzes the strength of shock-induced vortices due to these unsteady blade-row interactions. The time-accurate URANS code, TURBO, was used to generate periodic, quarter annulus simulations of the Blade Row Interaction compressor rig. Both time-averaged and time-accurate results compare well with experimentally-observed trends. It was observed that vortex shedding was synchronized to the passing of a rotor bow shock. Normal and large shock-induced vortices formed on the stator trailing edge immediately after the shock passing, but the large vortices were strengthened at the trailing edge due to a low-velocity region on the suction surface. This low velocity region was generated upstream of mid-chord on the suction surface from a shock-induced thickening of the boundary layer or separation bubble, due to the rotor bow shock reflecting off the stator trailing edge and propagating upstream. The circulation of the shock-induced vortices increased with shock strength (decreased axial spacing) and stator loading. Most design tools do not directly account for unsteady effects such as blade-row interactions, so a model is developed to help designers account for shock-induced vortex strength with varying shock strength and stator loading. An understanding of the unsteady interactions associated with blade loading and rotor shock strength in transonic stages will help compressor designers account for unsteady flow physics early in the design process.

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