Few aircraft engine manufacturers are able to consistently achieve high levels of performance reliability in newly manufactured engines. Much of the variation in performance reliability is due to the combined effect of tolerances of key engine components, including tip clearances of rotating components and flow areas in turbine nozzles. This research presents system analysis methods for determining the maximum possible tolerances of these key components that will allow a turbine engine to pass a number of specified performance constraints at a selected level of reliability. Through the combined use of a state-of-the-art engine performance code, component clearance loss models, and stochastic simulations, regions of feasible design space can be explored that allow for a pre-determined level of engine reliability. As expected, constraints such as spool speed and fuel consumption that are highly sensitive to certain component tolerances can significantly limit the feasible design space of the component in question. Discussed are methods for determining the bounds of any components feasible design space and for selecting the most economical combinations of component tolerances. Unique to this research is the method that determines the tolerances of engine components as a system while maintaining the geometric constraints of individual components. The methods presented in this work allow for any number of component tolerances to be varied or held fixed while providing solutions that satisfy all performance criteria. The algorithms presented in this research also allow for an individual specification of reliability on any number of performance parameters and geometric constraints. This work also serves as a foundation for an even larger algorithm that can include stochastic simulations and reliability prediction of an engine over its entire life cycle. By incorporating information such as time dependent performance data, known mission profiles, and the influence of maintenance into the component models, it would be possible to predict the reliability of an engine over time. Ultimately, a time-variant simulation such as this could help predict the timing and levels of maintenance that will maximize the life of an engine for a minimum cost. / Master of Science

The present study focuses on local heat transfer characteristics on the tip and near-tip regions of a turbine blade with a flat tip, tested under transonic conditions in a stationary, 2-D linear cascade consisting of seven blades, the three center blades having a variable tip clearance gap. The effects of tip clearance and exit Mach number on heat transfer distribution were investigated on the tip surface using a transient infrared thermography technique. In addition, thin film gages were used to study similar effects on the near-tip regions at 94% based on engine blade span of the pressure and suction sides. The experiments were conducted at the Virginia Tech transonic blow-down wind tunnel facility with a seven-blade linear cascade. Surface oil flow visualizations on the blade tip region were carried-out to shed some light on the leakage flow structure. Experiments were performed at three exit Mach numbers of 0.7, 0.85, and 1.05 for two different tip clearances of 0.9% and 1.8% based on engine blade span. The exit Mach numbers tested correspond to exit Reynolds numbers of 7.6 x 105, 9.0 x 105, and 1.1 x 106 based on blade true chord. The tests were performed with a freestream turbulence intensity of 12%. Results at 0.85 exit Mach showed that an increase in the tip gap clearance translates into a 12% increase in the heat transfer coefficients on the blade tip surface. Similarly, at 0.9% tip clearance, an increase in exit Mach number from 0.85 to 1.05 also led to a 24% increase in heat transfer on the tip. High heat transfer was obtained at the leading edge area of the blade tip, and an increase in the tip clearance gap and exit Mach number augmented this leading edge heat transfer. At 94% of engine blade span on the suction side near the tip, a peak in heat transfer was observed in all test cases at an s/C of 0.66 due to the onset of a downstream leakage vortex. At the design condition, this peak represents an increase of a factor of 2.5 from the immediate preceding s/C location. An increase in both the tip gap and exit Mach number resulted in an increase, followed by a decrease in the near-tip suction side heat transfer. On the near-tip pressure side, a slight increase in heat transfer was observed with increased tip gap and exit Mach number. In general, the suction side heat transfer is greater than the pressure side heat transfer as a result of the suction side leakage vortices. / Master of Science

<p>With increasingly tighter tip clearances in modern turbomachinery, it is essential to precisely measure this parameter during turbomachinery characterization. Benefits from measuring tip clearances include monitoring the structural integrity of the machine and estimating aerodynamic losses incurred due to leakage flows. At present, capacitance probes are one of the most commonly used sensors for tip clearance measurements in turbomachines as they are accurate and robust. One of the main challenges when using capacitance probes is properly calibrating the sensors, which usually involves complex positioning systems and blade representative targets. This manuscript describes in detail the development of a methodology for in-house calibration of capacitance probes for tip clearance measurements. A novel calibration procedure that does not involve rotating components is investigated and compared against established calibration methods. First, a calibration bench was developed to demonstrate the static and dynamic performance of the acquisition system and perform quasi-static as well as dynamic calibrations in a controlled environment. An in-situ methodology was then developed to calibrate the sensors once installed in a two-stage rotating turbine rig. The proposed methodology does not require complex positioning systems and a regression analysis using a least squares scheme resulted in a coefficient of determination of 0.9998. The calibration was validated using specially designed instrumentation at various speeds that span the operating envelope of the rig. A Bayesian model that was developed to estimate measurement uncertainties for each method showed that uncertainties as low as ± 5μm can be achieved with the proposed system. The proposed methodology was used in a two-stage turbine rig. Measurements taken at three different circumferential locations were subsequently used to map the spatial distribution of tip clearances throughout the speed operational envelope of the turbine. Finally, a reduced order rotor displacement model was developed and fitted to capacitance probes data. The work presented in this thesis lays the foundation for high fidelity tip clearance measurement capabilities at the Purdue Experimental Turbine Aerothermal Laboratory and can be implemented into any rotating rig. </p>

The objective of this research is to investigate the effects of tip clearance on the stage and component performance in a high-speed centrifugal compressor. The experimental data were compared against results from a numerical model to assess the ability of the numerical simulation to predict the effects of tip clearance. Experimental data were collected at Purdue University on the Single Stage Centrifugal Compressor (SSCC), a high-speed, high-pressure ratio test compressor sponsored by Honeywell Aerospace. Numerical simulations were completed using the ANSYS CFX software suite and part of the research computing clusters located at Purdue University.<div><br></div><div>Two tip clearances were tested, the nominal tip clearance and a tip clearance that is 66% larger than the nominal clearance, at speeds from 60% to 100% corrected speed. To compare data points with different tip clearances, various parameters were evaluated, and one was chosen. The value of TPR/inlet corrected mass flow rate best represented similar loading conditions, and thus similar incidences, for each tip clearance and was chosen as the best method for comparing similar data points taken with different clearances. Stage and component performance were focused on the sensitivity of each performance parameter to the changing of the tip clearance. The stage total pressure ratio and stage efficiency showed moderate sensitivity while the stage work factor showed much lower sensitivity. The impeller is more sensitive to changing tip clearances than the stage is, showing greater changes when comparing data from each tip clearance. The diffuser was on the same order of sensitivity as the impeller, with marginally higher sensitivities for some parameters. It was found that by the typical performance metrics, the diffuser performs worse at the nominal clearance than at the larger clearance. Upon further investigation though, the impeller is providing a higher static pressure and therefore, more diffusion, at the nominal clearance so the diffuser must perform less diffusion during nominal clearance operation.<br></div><div><br></div><div>To assess the validity of a prediction of the performance and sensitivity of the stage and components to the tip clearance, a numerical model was developed and validated. The numerical model was able to reasonably predict the stage performance with better comparisons of performance in the impeller and worse in the diffuser. The instrumentation in the experiment was replicated in the software to calculate performance the same way it is calculated experimentally so that the results would be comparable. While the performance of the stage and components was lacking in some areas, the trends predicted were similar to those calculated from the experimental data. As with the performance, the trends in the impeller matched very well between the experiment and the numerical simulation. The trends in stage and diffuser performance were predicted more accurately than the stage and diffuser performance maps and were able to capture the magnitude of the change in performance caused by changing the tip clearance. <br></div>

Recent design trends in steam turbines strive for high aerodynamic loading and high aspect ratio to meet the demand of higher efficiency. These design trends together with the low structural frequency in last stage steam turbines increase the susceptibility of the turbine blades to flutter. Flutter is the self-excited and self-sustained aeroelastic instability phenomenon, which can result in rapid growth of blade vibration amplitude and eventually blade failure in a short period of time unless adequately damped. To prevent the occurrences of flutter before the operation of new steam turbines, a compromise between aeroelastic stability and stage efficiency has to be made in the steam turbine design process. Due to the high uncertainty in present flutter prediction methods, engineers use large safety margins in predicting flutter which can rule out designs with higher efficiency. The ability to predict flutter more accurately will allow engineers to push the design envelope with greater confidence and possibly create more efficient steam turbines. The present work aims to investigate the influence of tip clearance flow on the prediction of steam turbine flutter characteristics. Tip clearance flow effect is one of the critical factors in flutter analysis for the majority of aerodynamic work is done near the blade tip. Analysis of the impact of tip clearance flow on steam turbine flutter characteristics is therefore needed to formulate a more accurate aeroelastic stability prediction method in the design phase.Besides the tip leakage vortex, the induced vortices in the tip clearance flow can also influence blade flutter characteristics. However, the spatial distribution of the induced vortices cannot be resolved by URANS method for the limitation of turbulence models. The Detached-Eddy Simulation (DES) calculation is thus applied on a realistic-scale last stage steam turbine model to analyze the structure of induced vortices in the tip region. The influence of the tip leakage vortex and the induced vortices on flutter prediction are analyzed separately. The KTH Steam Turbine Flutter Test Case is used in the flutter analysis as a typical realistic-scale last stage steam turbine model. The energy method based on 3D unsteady CFD calculation is applied in the flutter analysis. Two CFD solvers, an in-house code LUFT and a commercial software ANSYS CFX, are used in the flutter analysis as verification of each other. The influence of tip leakage vortex on the steam turbine flutter prediction is analyzed by comparing the aeroelastic stability of two models: one with the tip gap and the other without the tip gap. Comparison between the flutter characteristics predicted by URANS and DES approaches is analyzed to investigate the influence of the induced vortices on blade flutter characteristics. The multiple induced vortices and their relative rotation around the tip leakage vortex in the KTH Steam Turbine Flutter Test Case are resolved by DES but not by URANS simulations. Both tip leakage vortex and induced vortices have an influence on blade loading on the rear half of the suction side near the blade tip. The flutter analysis results suggest that the tip clearance flow has a significant influence on blade aerodynamic damping at the least stable interblade phase angle (IBPA), while its influence on the overall shape of the damping curve is minor. At the least stable IBPA, the tip leakage vortex shows a stabilization effect on rotor aeroelastic stabilities while the induced vortices show a destabilization effect on it. Meanwhile, a non-linear unsteady flow behavior is observed due to the streamwise motion of induced vortices during blade oscillation, which phenomenon is only resolved in DES results.

No description available.

In axial compressors with a relatively large blade tip clearance, an unsteady phenomenon denoted as rotating instability (RI) can be detected when the compressor is throttled to the operating points near the stability limit. In the frequency domain, RIs are shown as a hump lower than the blade passing frequency. This indicates an increase in noise level and might cause blade vibration and other undesirable structural issues. In this thesis, a comprehensive study on RIs is performed based on an axial compressor rotor row of the Low Speed Research Compressor at Technische Universität Dresden. Three blade tip clearances are investigated, and a groove casing treatment is mounted over the shroud for flow control. Methods of numerical modeling are evaluated, and zonal large eddy simulation is selected as the numerical model. By analyzing the flow properties and applying the dynamic mode decomposition, the coherent flow structure corresponding to the dominant frequency of RIs is extracted and visualized as the waves located in the blade tip region. The criteria for the appearance of RIs in the investigated research object are concluded.

Highly loaded single stage gas turbines are being developed to minimize the turbine size and weight. Such highly loaded turbines often result in transonic flows, which imply a reduction in the efficiency due to the shock losses. The efficiency of a turbine is defined as the ratio between the real work extracted by the turbine rotor from the fluid and the maximum available enthalpy for a given pressure ratio. The relationship between turbine performance and design parameters is not yet fully comprehended due to the complexity of the flow field and unsteady flow field interactions. Hence, experimental and numerical studies remain necessary to understand the flow behavior at different conditions to advance the state of the art of the prediction tools. The purpose of the current research is to develop a methodology to determine the efficiency with an accuracy better than 1 % in a cooled and uncooled high pressure (HP) turbine tested in a short duration facility with a running time of about 0.4s. Such low level of uncertainty requires the accurate evaluation of a large number of quantities simultaneously, namely the mass flow of the mainstream, the coolant, and leakage flows properties, the inlet total pressure and total temperature, the stage exit total pressure, the shaft power, the mechanical losses and the heat transfer. The experimental work is carried out in a compression tube facility that allows testing the turbine at the temperature ratios, Re and Mach numbers encountered in real engines. The stage mass flow is controlled by a variable sonic throat located downstream of the stage exit. Due to the absence of any brake, the turbine power is converted into rotor acceleration. The accurate measurement of this acceleration as well as those of the inertia and the rotational speed provides the shaft power. The inertia of the whole rotating assembly was accurately determined by accelerating and decelerating the shaft with a known energy. The mass-flow is derived from the measured turbine inlet total pressure and the vane sonic throat. The turbine sonic throat was evaluated based on a zero-dimensional model of the turbine. The efficiencies of two transonic turbines are measured at design and off-design conditions. The turbine design efficiency is obtained as 91.8 %. The repeatability of the measurements for 95% confidence level varies between 0.3 % and 1.1 % of the efficiency depending on the test case. The theoretical uncertainty level of 1.2 % is mainly affected by the uncertainty of exit total pressure measurements. Additionally, the effect of vane trailing edge shock formations and their interactions with the rotor blade are analyzed based on the experimental data, the numerical tools and the loss correlations. The changes of blade and vane performances are measured at mid-span for three different pressure ratios which influence the vane and rotor shock mechanisms. Moreover, the unsteady forces on the rotor blades and the rotor disk were calculated by integration of the unsteady static pressure field on the rotor surface.

A thermal tip clearance control system provides a robust and flexible means of manipulating the closure between the casing and the rotating blade tips in a jet engine, reducing undesirable tip leakage flows. This may be achieved using an impingement cooling scheme on the external casing of the engine in conjunction with careful thermal management of internal over-tip seal segment cavity. For a reduction in thrust specific fuel consumption, the mass flow rate of air used for cooling must be minimised, be at as low a pressure as possible and delivered through a light weight structure surrounding the rotating components in the turbine. This thesis first characterises the effectiveness of a range of external impingement cooling arrangements in typical engine casing closure system. The effects of jet-to-jet pitch, number of jets, inline and staggered alignment of jets, arrays of jets on flange, on an engine representative casing geometry are assessed through comparison of the convective heat transfer coefficient distributions in a series of numerical studies. A baseline case is validated experimentally. The validation data allowed the suitability of different turbulence closure models to be assessed using a commercial RANS solver. Importantly for each configuration the thermal contraction of an idealised engine casing is predicted using thermo-mechanical finite element models, at a series of operating conditions representing engine idle to maximum take-off conditions. Cooling is provided by manifolds attached to the outside of the engine. The assembly tolerance of these components leads to variation in the standoff distance between the manifold and the casing. For cooling arrangements with promising performance, the study is extended to characterise the variation in closure with standoff distance. It is shown that where a sparse array of non-interacting jets is used the system can be made tolerant of large build misalignments. The casing geometry itself contributes to the thermal response of the system, and, in an additional study, the effect of casing thickness and circumferential thermal control flanges are investigated. Restriction of the passage of heat into the flanges was seen to be dramatically change their effectiveness and slight necking of the flanges at their root was shown to improve the performance disproportionally. High temperature secondary air flowing past the internal face of the engine casing tends to heat the casing, causing it to grow. Experimental and numerical characterisation of a heat transfer within a typical over-tip segment cavity heat transfer is presented in this thesis for the first time. A simplified modelling strategy is proposed for casing and a means to reduce the casing heat pickup by up to 25 % was identified. The overall validity of the modelling approach used is difficult to validate in the engine environment, however limited data from a test engine temperature survey became available during the course of the research. By modelling this engine tip clearance control system it was shown that good agreement to the temperature distribution in the engine casing could be achieved where full surface external heat transfer coefficient boundary conditions were available.

The quest for higher performances and durability of modern aero-engines requires the understanding of the complex aero-thermal flow experienced in a multi-row environment. In particular, the high and low pressure turbine components have a great impact into the engine overall performance, and improvements in the turbine efficiencies can only be achieved through detailed research on the three-dimensional unsteady aerodynamics and heat transfer. The present thesis presents an experimental study of the aerothermodynamics in one and a half turbine stage, focusing on: the aero-thermal flow in the overtip region of a transonic highly loaded high pressure (HP) rotor, and the aerodynamics and heat transfer of an innovative low pressure (LP) stator with a multi-profile configuration placed downstream of the high pressure turbine, within an s-shaped duct. Advanced instrumentation and measurement techniques were used and developed to perform the experimental investigation in a short-duration turbine test rig where both high spatial and time accuracy is indispensable. The flow field at the rotor shroud was investigated with simultaneous measurements of heat transfer, static pressure and blade tip clearance by using fast response pressure, wall temperature and capacitance probes. Through repeat experiments at the same turbine operating point, the time-averaged and time-resolved adiabatic wall temperature and convective heat transfer coefficient were evaluated. In the frame of new engine architectures, a novel stator for an LP turbine is proposed with a multi-splitter layout that represents a new design solution towards compact, lighter and performing aero-engine turbomachinery. It contains small aero-vanes and large structural aerodynamic airfoils which are used to support the engine shaft and house service devices. The research focuses on the experimental investigation of the global performance, aerodynamics and thermodynamics of this novel HP-LP vane layout. The turbine was / Lavagnoli, S. (2012). On the Aerothermal Flow Field in a Transonic HP Turbine Stage with a Multi-Profile LP Stator Vane [Tesis doctoral no publicada]. Universitat Politècnica de València. https://doi.org/10.4995/Thesis/10251/17799 / Palancia