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At present, both governmental and private aviation enterprises are trying to develop aircraft that are swifter, more lightweight, and more cost-effective in order to maintain competitiveness on both the domestic and global stage. Continuous innovation and sustainability efforts are necessary to achieve advancements in aviation systems, such as fans, compressors, combustors, and turbines, as well as sub-systems, like engine seals. Advanced engine seals exhibit considerable potential in enhancing the engine's pressure ratio and cycle temperatures, leading to reduced engine weight, increased thrust, and improved fuel economy. In this study, a novel multistage seal idea has been proposed for the supercritical carbon dioxide (sCO2) heat recovery unit for aircraft engines that utilizes the established Elastohydrodynamic (EHD) lubrication theory. This study utilized a physics-based computer simulation modeling technique. Because of the complex and non-linear interactions between the seal and the fluid domain in Fluid-Solid-Interactions (FSI), a design approach that combines Finite Element Analysis (FEA) and Computational Fluid Dynamics (CFD) was used. Utilizing the latest advancements in commercially accessible software, the COMSOL Multiphysics program was utilized to conduct fully linked FSI simulations. The modeling technique was comprehensively explained, and the findings were fully examined. The leakage was found to be 0.016 kg/s at 19 MPa. The suggested seal system has the potential to serve as a substitute for the current sealing technologies available in the market.

In recent studies, a turbine film cooling shaped hole designed by adjoint optimization techniques (X-AOpt) was found to have substantially increased film cooling performance. This thesis investigated the X-AOpt film cooling hole under varying operating conditions, to include experimentation with geometry variations of the hole itself and integration into a multirow configuration to model the pressure side of a turbine airfoil. Two aspects of the X-AOpt hole's geometry contribute to its improved cooling effectiveness: the shape of the internal geometry (which improves the diffuser performance) and the external protrusions (which generate counter rotating vortices that laterally spread the coolant more effectively over the surface of the model airfoil). Results from experimental evaluations of the X-AOpt hole with the external protrusions removed showed that 40\% of the improvement in film cooling effectiveness is due to the internal geometry, and 60\% of the improvement is due to the external protrusions. Variations in lateral spacing, or pitch, of the X-AOpt holes provided further insight into the optimized interaction of the counter rotating vortices generated by the hole's external protrusions. The X-AOpt hole was then integrated into a multirow configuration, where it was evaluated in both adiabatic and overall cooling effectiveness experiments. The improvements in film cooling effectiveness from the implementation of the X-AOpt hole were characterized, and a superposition analysis was completed to compare actual performance to predicted performance.

NASA Glenn Research Center has laid out goals to meet the needs of modern day commercial air travel. The goals laid out by NASA aim to reduce the fuel consumption while improving upon the performance parameters of future aircraft. One approach that has been researched to meet these standards is the implementation of a boundary layer ingesting (BLI) inlet. It has been shown that BLI inlet designs can achieve much of the targets set out by NASA. However, BLI inlets can harm the performance of the engine due to the large boundary layer growth which develops at the bottom surface of the inlet. Later studies have shown that the placement of a fan at the aerodynamic interface plane can attenuate the losses associated with the boundary layer growth. However, the severe changes in total pressure and high incidence angles have shown to hinder the performance of the fan and provided risk for the fan to fail due to stall conditions. The goal of this study is to optimize the design of the fan to survive the harsh conditions of the BLI inlet. The optimization was broken up into three primary phases. Phase 1 utilized a 2D optimization approach which sought to independently optimize cross-sections of a single passage under tangentially averaged inlet flow profile, and then restacked the optimized cross-sections to form a new rotor geometry. Phase 2 built upon the optimized designs of the first phase and optimized a single passage of the rotor modeled with a tangentially averaged inlet flow profile using a high end computational fluid dynamics tool called TURBO. A toolbox, labeled the Blade Altering Toolbox, was created for the second phase in order to manipulate the shape and design of the 3D rotor. Phases 1 and 2 were both used to produce a narrowed down design in order to save computational costs. Models of the Phase 2 optimized geometries were analyzed as both a single passage model using a tangentially averaged inlet profile which contained information of the distortion field, and a full annulus model containing the entire distortion profile prescribed at the aerodynamic interface plane. The performance variables of the single passage model compared to their respective full annulus models were virtually identical which suggests that the rotor can be accurately optimized using a single passage model with a tangentially averaged inlet profile that contains the distortion field. Phase 3 then further optimized the designs from Phase 2 by modeling the full annulus with the full distortion field prescribed at the aerodynamic interface plane. The optimized designs each showed a significant reduction in the blade separation and have mitigated the risks of stall occurring within the distortion field. As a result, the optimized designs have also demonstrated a significant improvement in the adiabatic efficiency with a small reduction of the total pressure ratio.

Turbomachinery components, such as the low pressure turbine, are highly complex rotating machines, therefore, conducting fundamental fluid mechanics studies in them is exceedingly difficult. For this reason, testing is generally completed in facilities such as linear cascades, like the one present in the Low Speed Wind Tunnel Facility at AFRL, which typically utilize a low freestream turbulence intensity, when in reality, the freestream turbulence intensity in a full, rotating low pressure turbine is likely much higher. Slightly elevating the freestream turbulence intensity (e.g., 3%) typically improves the Reynolds-lapse characteristics of a blade profile by affecting the transition process, reducing the detrimental effects of laminar boundary layer separation, and shifting the knee in the loss curve. Front loaded blades are more resistant to separation, however, they can experience high losses in the endwall region due to the complex vortical structures present. Therefore, a better understanding whether high levels of freestream turbulence intensity will increase the overall losses generated in the passage is important. An intial study with a jet based active grid was completed on the L2F blade. Based of the insight gained from that study, a new mechanical agitator based active grid was implemented into a linear cascade of L3FHW-LS blades in order to more effectively study how elevated FSTI impacts the endwall flow behavior and loss production. Coefficient of pressure measurements, three planes of SPIV, two additional planes of flow visualization, and three planes of total pressure loss measurements were collected. Impacts of incoming turbulence on the endwall losses as well as the endwall flow structures were assessed.

Integrally bladed disks, or blisks, are critical components within compressors of modern turbomachinery, and during operation can exhibit complex dynamic behaviors. Understanding and accurately modeling these dynamics is crucial for blisk design and safe operation. To model blisk dynamics, large finite-element models containing millions of degrees of freedom are often used. However, simulating blisk dynamics using these models can be computationally expensive and frequently infeasible. As such, physics-based reduced-order models are currently the preferred method for modeling blisk behavior and performing stress analyses. These models typically are used to run probabilistic analyses such as Monte Carlo simulations for capturing the effects of variabilities in blisk dynamics produced from unavoidable deviations in material properties and geometry, called mistuning. Mistuning is known to cause energy localization within the blisk, resulting in amplified response amplitudes and greater stresses than those present in a nominally cyclic-symmetric blisk. Because the uncertainties introduced by mistuning cannot be removed, it is of great interest to both understand the effects of mistuning and identify it in order to accurately predict forced responses during operation when designing a blisk. Additionally, to mitigate large response amplitudes, nonlinear damping mechanisms such as ring dampers or shrouds can be employed, which dissipate energy through friction contacts. For both linear and nonlinear systems, physics-based reduced-order models, however, often cannot incorporate experimental data to improve their accuracy and applicability to as-manufactured models and/or rely on experimental modal identification that can be challenging within the often unavoidable high-modal density regimes of blisk structures. To address these challenges, this work presents novel data-driven methods for 1) forced-response prediction of mistuned blisks with and without introducing friction nonlinearities, and 2) identifying the mistuning of as-manufactured blisks. All presented methodologies focus on sector-level, frequency-domain formulations to reduce necessary training data and avoid difficulties associated with time-domain data generation and experimental collection. First, a data-driven method for forced-response prediction is presented and demonstrated for a lumped mass model mimicking a mistuned blisk. This method is based on two artificial neural networks and a cyclic coupling procedure, and is applicable to linear mistuned cyclic structures in general. Next, a novel data-driven approach for mistuning identification is developed. Unlike previous methods, this approach does not require any modal analyses nor isolation of individual blades, and it removes all effects of forcing amplitude via a developed robust data processing, selection, and conditioning procedure. This approach is validated for a high-dimensional finite-element blisk model considering significant measurement noise similar to that observed experimentally. A first-of-its-kind physics-informed machine learning modeling approach for forced response prediction is developed that incorporates physical laws directly into a network architecture, still maintains a sector-level viewpoint, and only considers response data from easily experimentally measurable points along blade tips. With this approach, a series of numerical techniques are developed to significantly improve robustness and generalizability. Lastly, to predict forced responses of blisk systems containing contact nonlinearities, the original linear approach is extended using a harmonic-balance formulation. This includes the development of an efficient data selection and training scheme to further reduce the amount of necessary training data. Validation is shown for a nonlinear lumped mass model representative of a mistuned blisk with a friction ring damper and significant mistuning and nonlinear effects.

Improving the cooling effectiveness of turbine blades is one of the key drivers in gas turbine engine performance. Transpiration cooling describes an ideal arrangement that maximises this cool-ing effectiveness through the use of porous structures. High internal surface area and uniform cool-ant distribution are some of its main benefits. However, physical limitations in manufacturing and materials mean that this arrangement is impractical to implement. This research identifies some of the key properties that makes transpiration cooling success-ful and recreate them with geometric features already within the bounds of existing manufacturing capabilities. The use of simple, small round film cooling holes, in tight staggered arrays at varying locations around the blade profile was found to generate exceptional levels of film effectiveness. A series of experimental campaigns were carried out to quantify the effects of different parameters in driving this high film effectiveness. Features placed on the early suction surface of the blade running at low blowing ratios were seen to be perform particularly well, with film effectiveness levels above 90% being maintained to the trailing edge. Analytical modelling of the results demonstrated that the overall array of film holes generates films that behave similarly to tangential injection slots. A saturated film core was identified, which remained effectively separate from the turbulent boundary layer. Further analysis suggested that the performance of the suction surface films could be explained by these operating in a laminar-like regime, contradicting many commonly held assumptions used for suction surface film modelling. Investigations into the nature of the interactions between the individual film cooling jets were also carried out using numerical simulations. These illustrated some of the beneficial outcomes of positioning film holes in close proximity to each other. The effects of counter rotating vortex pairs and boundary layer velocity profiles were identified as key mechanisms driving these interactions. Crucially, the magnitude of these interactions was found to outweigh any effect of individual film hole geometry. The present work culminated in a series of metal effectiveness tests on turbine blade geometries. These were used to obtain a complete measure of the performance of high porosity cooling arrays within a double walled cooling scheme. The aim of these experiments was to showcase the potential breakthrough in film cooling that had been unlocked by the high porosity array approach. Results in this experiment demonstrated that metal effectiveness values for the proposed systems already exceeded those of mature state of the art cooling arrangements.

Gas turbines are essential in electrical power generation and aircraft applications. One way to increase the efficiency of a gas turbine engine is to increase the combustor exit temperature. However, temperatures higher than the melting point of turbines located downstream can result in serious thermal failures. Therefore, these high temperatures create a need to design aggressive cooling schemes for engine sections to prevent component damage. However, owing to the complexity of the flow in the engine, it is essential to understand how coolant flows interact with engine passage flow. This thesis discusses three experimental studies relating to the cooling of gas turbines:(1) Evolution of secondary flows: Due to the geometry of turbine vanes, various undesired flows are developed in the vane passage, which are termed as secondary flows. The flowrates of the injected coolants affect the way in which these secondary flows are generated and transported. A detailed description of the vane passage secondary flowfield for a variety of coolant flowrates is provided and discussed. (2) Film cooling performance: The injected coolant forms a film on the surfaces to be cooled to protect them from failure. The cooling performance of a novel coolant injection scheme is reported in this study. Coolant transport is recorded using temperature and velocity measurements. The cooling performance on the vane passage surfaces is discussed using these transport measurements and compared with injection scheme currently used in the engine. (3) Decay of turbulence: Vane geometry leads to changes in turbulence features of the flow, which are known to affect the cooling of the vane surfaces. Such changes to turbulence were measured and discussed. Numerical simulations using Reynolds-averaged Navier-Stokes turbulence models were also performed for the same vane geometry. A comparison between the computed and the measured parameters is also presented. The results of these studies are meant to help gas turbine designers in reducing the amount of required coolant which would lead to an increase in the gas turbine efficiency.

Due to regulatory and socio-economic drivers, there is an ongoing process of transforming the energy market. This process aims to reduce emissions and improve efficiency, thus reducing costs and subsequently increasing profit in the power generation sector and the energy market as a whole. The main goal has been the transition to renewable energy technologies, reducing carbon-emitting intense traditional sources. As such, power generation from natural gas is expected to reduce in the mid-term future, albeit, at a slow pace due to benefits associated with operational flexibility, efficiency and low carbon footprint with respect to alternatives. Moreover, turbomachinery components are employed across all industrial sectors and for this reason there is still a profound need to improve their performance. Experimental data-sets are needed to verify the numerical tools, which are currently the backbone in turbomachinery research and are steadily becoming more sophisticated even in initial design space exploration studies. Due to the vast improvements in turbomachinery component efficiencies over the past century, nowadays marginal efficiency improvements are more and more difficult to attain and also detect. For these reasons, high-fidelity experimental validation is as relevant as ever. The global objective of the present work is to promote the use of rigorous experimental methods regarding fast response aerodynamic probe measurements in the challenging flow-field encountered in turbomachinery representative configurations. The present work was carried out within the scope of an academic research project aiming to advance the state-of-the-art for point-wise pressure-based multi-sensor fast response aerodynamic probes (FRAP). Miniature four-sensor FRAP have been developed, the smallest of which features an external diameter of 3 mm. These prototype FRAP are additively manufactured (AM) from stainless steel using a binder jetting technique, thus allowing the fast and cost-effective development of various FRAP which are used for further investigations in the context of this work. The FRAP-4S includes silicon piezo-resistive pressure sensors which are soft-bonded with a silicon elastomer and are housed directly beneath the sensing hole. The resulting measurement bandwidth is in the order of 45 kHz without any numerical compensation. At first, the development and application of new instrumentation able to perform three-dimensional time-resolved measurements are described. Details on the design, packaging, assembly and calibration processes of the newly developed instruments are given, highlighting their advantages and limitations over traditional instrumentation used in turbomachinery testing. Moreover, post-processing and uncertainty quantification (UQ) methods are developed, tailored to aerodynamic probes. A work-flow to quantify the uncertainties related to the derived flow quantities is presented. The breakdown and contributions of the elemental uncertainty sources is provided both for pneumatic and FRAP. Concerning FRAP, a method to quantify the temperature dependent pressure hysteresis is described and is subsequently included in the uncertainty calculation. Additionally, the temporal drift of the piezoresistive pressure sensors is included in the UQ method and it is shown that its contribution varies between 30% to 75% and thus should always be included in the UQ of FRAP. A novel iterative data-reduction algorithm is developed to account for the compressibility sensitivity demonstrated from aerodynamic probes and a method to quantify the convergence criteria of the process is thoroughly described. The algorithm is focused on correcting the systematic errors in the time-varying flow-quantities when improper calibration files are used to evaluate the unsteady FRAP data. The methodology is used to post-process the experimental results obtained in a one-and-a-half stage axial turbine research facility. The results indicated differences for mainly static pressure and pitch angle, which vary primarily with the magnitude of Mach number fluctuations and secondarily with the time-averaged Mach number. The fact that the errors appear in the ensemble-averaged results indicates that these are significant and are not averaged out. A unique experimental dataset is obtained studying the effect of the FRAP-4S size on the measured turbulence levels (Tu) and related turbulence information. The mean Tu, measured by all FRAP-4S, match levels recorded by hot-wire anemometry studies in similar turbine configurations from the literature. With regards to probe size, small differences were found in mean and time-varying Tu both behind a vane and rotor blade rows with these quantities being proportional to the FRAP size. Statistically significant differences in the average Tu were found at the location of the rotor tip leakage flow where, surprisingly, the larger probes indicated lower turbulence intensity. Measured length scales and the degree of turbulence anisotropy were more significantly affected by the probe size. The most stark differences were found on the spatial distribution of the integral length scales at first stator exit. A dedicated spectral filtering approach is proposed to improve the estimation of the length scales and is shown to perform well, yet further refinement is needed to improve its reliability. Finally, an extensive investigation is performed on methodologies used to derive turbulence information from unsteady pressure measurements. The applicability and limitations of these methods are discussed and strategies to correct the turbulence levels is demonstrated. Thus, corrected turbulence intensities are within a few percent (< 2%) difference with respect to the true value. This is substantial improvement considering that these methods are providing non-physical results at particular flow regimes within the turbine.

Supercritical CO2 (sCO2) power cycles are next generation power technologies offering improved efficiency, lower customer costs, and less water usage. They hold significant potential in many different industries such as nuclear power production, ship propulsion, geothermal power, solar power, and fossil fuel power plants. One of the largest hurdles preventing this milestone is the lack of suitable shaft seals for sCO2 operating conditions. This component issue needs to be addressed for the next generation nuclear turbine and compressor development to progress. In this study, a novel Elastohydrodynamic (EHD) seal is proposed. This proposed design is capable of handling high pressure and high temperature while being a scalable shaft seal for sCO2 turbomachinery. The potential benefit of this design includes low leakage, minimal wear, low cost, and no stress concentration. The focus of this study was to perform a proof-of-concept study utilizing physics-based computer simulations. A Fluid-Structure-Interaction (FSI) modeling approach was utilized, and the simulations were carried out in COMSOL Multiphysics software. The results indicated that the proof-of-concept study was successfully demonstrated. The leakage rate obtained from the physics-based computer simulations followed a quadratic form with a peak at m = 0.075 kg/s at Pin = 15 MPa and then decaying to less than m = 0.040 kg/s at Pin = 30 MPa. This trend signifies that the design has the potential to be optimized and be tailored to be a solution to the shaft sealing issues found during the operation of sCO2 turbomachinery.

Supercritical CO2 (sCO2) power cycles are superior to traditional water-based, air-breathing, direct-fired, open Brayton cycles; or indirect-fired, closed Rankine cycles in terms of efficiency and equipment footprint. They hold great potential in nuclear power production, fossil fuel power plants, concentrated solar power, geothermal power, and ship propulsion. To unlock the potential of sCO2 power cycles, technology readiness must be demonstrated on the scale of 10 – 600 MWe and at the sCO2 temperatures and pressures of 350 – 700 ºC and 15 – 30 MPa for nuclear industries. The lack of suitable shaft seals at sCO2 operating conditions is one of the main challenges at the component level. So far, conventional seals are incapable of handling sCO2 pressure and temperature in one way or another. This study proposes a patented novel Elasto-Hydrodynamic (EHD) high-pressure, high temperature, and scalable shaft seal for sCO2 turbomachinery that offers low leakage, minimal wear, low cost, and no stress concentration. The focus of this paper was to conduct a proof-of-concept study with the help of computer simulations. To this end, a fully coupled Fluid-Structure Interaction (FSI) modeling approach was adopted, and the simulations were carried out in COMSOL Multiphysics software. The modeling approach was presented thoroughly, the results were discussed, and the subsequent research steps were highlighted. Under the EHD mechanism: the higher the pressures are, the tighter the sealing becomes while sustaining a continuous sCO2 film. a 2" diameter shaft static test rig was designed at Georgia Southern University before the actual dynamic testing at Sandia National Laboratories for proof-of-concept purposes. For simplicity, the working fluid was chosen to be Nitrogen, and the tests were conducted at room temperature. The test rig consisted of a 16.5 MPA N2 tank, steel tubing with compression type fittings, and OMEGA – PX5500C0-2.5KA10E pressure sensor to measure the pressure of the N2 after the pressure regulator from the tank, a chamber housing the 2" steel shaft and EHD seal (two prototype steel seals with seal clearances of 25 μm and 50 μm), a CONAX – K-INC12-U-T3(5FT)-7.5" temperature sensor to collect temperature data in the chamber, and an OMEGA – FMA-1623AI mass flow meter with temperature and pressure measurement capability to evaluate conditions after the test chamber. The pressure, temperature, and mass flow rate data were collected via a National Instruments DAQ – Module NI-9205 and a LabVIEW program. The tests were conducted for an inlet pressure of 0.1 ~ 15 MPa and at two different clearances of 25 μm and 50 μm. The mass flow rate exhibited a quadratic trend with increased pressure, which proved the hypothesis.

Turbomachinery, like jet engines and industrial gas turbines in power plants, are very advanced and complex machines. Due to the complexity and cost of modern turbomachinery, there is active research in accurately predicting the physical system dynamics using computational models. Two big mechanisms that affect the structural response are the prestress effects from high rotational speeds and mistuning effects from tolerance deviations, wear, or damage. Understanding the role these two mechanisms play in the computational modeling of these systems is an important step toward a complete digital twin of an entire jet engine. There previously existed modeling methods that enabled each to be analyzed independently, but not simultaneously in an efficient manner. This will be one of the focus points of this dissertation. The other focus being an experimental investigation into exciting system resonances of a rotating bladed disk using air jets. These experiments will be used to validate the computational modeling method developed.This dissertation has three primary objectives. The first objective is to present reduced order modeling methods that allow for the efficient modeling of coupled systems and rotating systems, both with small or large mistuning. By efficiently including these mechanisms, more realistic boundary conditions can be used to help validate the reduced order models (ROMs) with experimental data. Both modeling methods create models a fraction of the size of the full model while retaining key dynamic characteristics of the full model. The second objective of this work is to show the capability of air jets in exciting synchronous and non-synchronous vibrations in a rotating bladed disk. Much previous research in this field focused on experiments with stationary systems. These tests can help isolate specific mechanisms that may be present in bladed disks, but may limit the applicability of the results to actual rotating systems. This work presents a method of using air jet excitation to cause the desired vibratory response on a rotating bladed disk. The blade vibrations are measured using a non-contacting stress measurement system; the blade tip timing system used in this work uses lasers to measure when a blade arrives at a specific circumferential location and can use the deviations in arrival times to calculate the blade tip deflection amplitudes. This allows the system to rotate and does not change the structural properties of the blades themselves. The third and final objective of this work is to use the presented ROM methods to match the experimental results and investigate the deviations in the results of the models and experiments. The insights into updating computational models such that they predict physical system responses will help inform areas that need more attention in future modeling methods.

This document presents the development and implementation of a new generation of double-sided heat-flux gauges at The Ohio State University Gas Turbine Laboratory (GTL) along with heat transfer measurements for film-cooled airfoils in a single-stage high-pressure transonic turbine operating at design corrected conditions. Double-sided heat flux gauges are a critical part of turbine cooling studies, and the new generation improves upon the durability and stability of previous designs, while also introducing high-density layouts that provide better spatial resolution. These new customizable high-density double-sided heat flux gauges allow for multiple heat transfer measurements in a small geometric area such as immediately downstream of a row of cooling holes on an airfoil. Two high-density designs are utilized: Type A consists of 9 gauges laid out within a 5 mm by 2.6 mm (0.20 inch by 0.10 inch) area on the pressure surface of an airfoil, and Type B consists of 7 gauges located at points of predicted interest on the suction surface. At the same time, improvements to the manufacturing and installation processes for single gauges increased the survival rate of the gauges from 47% to 84%.Both individual and high-density heat flux gauges are installed on the blades of a transonic turbine experiment for the second build of the High-Pressure Turbine Innovative Cooling program (HPTIC2). Run in a short duration facility, the single-stage high-pressure turbine operated at design-corrected conditions (matching corrected speed, flow function, and pressure ratio) with forward and aft purge flow and film-cooled blades. Gauges are placed at repeated locations across different cooling schemes in a rainbow rotor configuration. Airfoil film-cooling schemes include round, fan, and advanced shaped cooling holes in addition to uncooled airfoils. Both the pressure and suction surfaces of the airfoils are instrumented at multiple wetted distance locations and percent spans from roughly 10% to 90%. Results from these tests are presented as both time-average values and time-accurate ensemble averages in order to capture unsteady motion and heat transfer distribution created by strong secondary flows and cooling flows.Temperature and Stanton number measurements from a baseline high temperature run with nominal cooling conditions are presented. A comparative analysis of time-average Stanton number across the various cooling schemes shows that no single cooling hole shape performs universally better than its counterparts. Rather, cooling benefits seen through a reduction of Stanton number vary among cooling schemes with changing span and wetted distance. This local variation of cooling benefits is even seen downstream of a single row of holes. Time-average heat transfer is presented for a Type A high-density gauge downstream of a row of fan-shaped cooling holes on the airfoil pressure surface. Results show Stanton number highest near the cooling holes and decreasing with increasing distance from the holes. This trend is explained by the radial migration of cooling flow as it exits the cooling hole. The gauge closest to the cooling hole is therefore exposed to the hot mainstream gas, imparting a large amount of heat to this location on the airfoil surface. A time-average temperature analysis at various experimental time windows gives confidence to the heat flux gauge results. A time-accurate Stanton number analysis is subsequently presented to paint a complete picture of heat transfer at this location on the airfoil pressure surface. Fluctuations in Stanton number reveal secondary peaks in the side-band gauges as the rotor passes through the low temperature vane wakes. This behavior suggests cooling flow unsteadiness as the cooling tracks move to uncover the side-band gauges. The shift in cooling track exposes them to the higher temperature mainstream gas and results in an increase in Stanton number.

A Jetcat P400 turbojet was used as a baseline benchmark to design a compressor stage which is capable of increasing mass flow by 50% without growing in size, in an effort to increase thrust to weight ratio. An axisymmetric through-flow design script was developed to create a two-stage mixed flow compressor design. Parametrization was utilized with a Single Objective Function Genetic Algorithm and an axisymmetric solver to optimize the flow-path and work input for efficiency. The design was materialized through GTSL developed tools T-Blade3 and geomturbo to create a baseline geometry. The baseline geometry was analyzed and improved upon with a Multi Objective Function Genetic Algorithm optimizing for mass flow and efficiency. Design parameters controlling blade geometry were defined and the genetic algorithm was paired to a 3D CFD meshing and analysis tool (FINE/TURBO). The first of two rotor-stator stages in the design is completed through aerodynamic evaluation. A 50% increase in mass-flow was achieved while maintaining the original size envelope and a 2:1 pressure ratio. The isentropic efficiency of the design is 80.667% at the design point and 81.79% peak efficiency at lower mass flow and pressure ratio.

The preliminary design of multistage axial compressors in gas turbine engines is typically accomplished with mean-line methods. These methods, which rely on empirical correlations, estimate compressor performance well near the design point, but may become less reliable off-design. For land-based applications of gas turbine engines, off-design performance estimates are becoming increasingly important, as turbine plant operators desire peaking or load-following capabilities and hot-day operability. The current work implements a one-dimensional stage stacking procedure, including a new blockage term, which is used to estimate off-design compressor performance and operability range of a 13-stage axial compressor used for power generation. The procedure utilizes stage characteristics which are constructed from computational fluid dynamics (CFD) simulations of groups of stages. The stage stacking estimates match well with CFD results. These CFD results are used to assess a metric which estimates the stall limiting stages.

One-way fluid structure interaction (FSI) is explored in two independent scenarios. First, a rectangular single expansion ramp nozzle (SERN) is analyzed through CFD. The SERN was simulated at three conditions: over-expanded, perfectly-expanded, and under-expanded (NPR 3.0, 3.671, and 4.75, respectively). A Schlieren image comparison between CFD results and experiment revealed that there were discrepancies in the location of the shock emanating from the nozzle throat and reflecting through the divergent ramp section. Structural analysis showed that the top plate deforms under load, deflecting outward from the flow by 31 micrometers when over-expanded and nearly 0.3 millimeters when under-expanded. Modal analysis was performed for unloaded, over- and under-expanded conditions, with the first several modes in good agreement with experimental results at the over-expanded condition. For the first four modes found with FEA, all were within 0.7% relative error, except the second mode that was 4.5% different.The second scenario in which FSI is applied is analysis and design of a rotor in a three-row stage with an inlet guide vane (IGV) and outlet guide vane (OGV) of a boundary layer ingesting (BLI) propulsor. The blade shape of the rotor was altered by adjusting the spanwise parameters of thickness distribution, sweep, lean, and chord multiplier in order for the design to satisfy structural requirements. Pressure loads were extracted from CFD and applied to the rotor blade surface for a thorough structural and modal analysis to gather an understanding of how the blade behaved under load. At design speed, the rotor achieved a safety factor of 1.6911 and deformed 0.88 millimeters radially and 5.05 millimeters tangentially. Several modes present possibilities of resonance from perturbations from both IGV and OGV. Although these conditions exist, tuning blade modes to avoid resonance is left to future work.

The first stage of an axial compressor was designed and optimized using 1D results from 3D CFD analysis. The compressor was based off of a small commercial jet engine compressor. Structural and modal analysis were performed on the compressor to ensure safety during testing. Modal analysis results were compared to Scanning Laser Vibrometry results. A custom test rig was designed and built utilizing a COTS turbocharger. Additive manufacturing was used to produce custom parts for the test rig, including the compressor test articles. Measurements of the flow were taken to find the performance characteristics of the turbine and custom compressor. The data was acquired using Labview and various COTS measurement devices. Issues with minimum thickness and high centrifugal loads led to multiple design changes which limited the expected performance of the compressor. Complications with additive manufacturing may have also reduced the performance of the compressor. Experimentation was performed at low speeds first. The test rig failed at around 60% design speed, likely due to issues with the custom compressor mounting plate. Performance characteristics were calculated with the limited data and speed lines were generated. The data was compared to the results from CFD analysis. The optimization system was shown to be useful in increasing the performance of a compressor stage. It was concluded that a low cost custom test rig using a turbocharger is a viable solution for small compressor testing. Additive manufacturing was shown to be useful in producing one-off custom designs, though is not a perfect solution. The overall process is a useful process for rapid development and testing of custom compressor designs for small engine applications.

This thesis covers the intensive research effort to elucidate the role of elevated temperature in deposition. Several experimental campaigns were conducted in this pursuit. The testing explored high temperature deposition with 0-10 micron Arizona Road Dust (ARD) with the intent of creating a yield strength model that included temperature effects and could be incorporated into the existing OSU deposition model.Experimental work was first conducted in the impulse kiln facility where small amounts of the test dust were placed on ceramic targets and rapidly exposed to temperatures between 1200K and 1500K. Trends in the packing factor confirmed the existence of two threshold values (1350K and 1425K) that could be linked to strength characteristics of the dust when exposed to high temperatures. Using the information obtained from the kiln experiments, HTDF testing was conducted between 1325K and 1525K. Exit temperatures were set at 25K intervals in this region with a constant jet velocity of 150 m/s. The capture efficiency data showed this trend with temperature and indicated a softening temperature and melting temperature of 1362K and 1512K respectively. With these critical values in hand, the Ohio State University Molten Model was created to modify yield strength with particle velocity and temperature. The model was tested using CFD and showed a good capability for capturing particle temperature effects in deposition from an impinging particle-laden jet.A subsequent test campaign was conducted to explore the effect of varying surface temperature on deposition. Hastelloy coupons with Thermal Barrier Coatings (TBCs) were subjected to a constant jet at 1600K jet and 200 m/s while being cooled via a backside impingement jet. Surface temperatures between 1455K and 1125K were impacted with 0-10 micron ARD while an IR camera monitored the surface. Coupons with higher coolant flowrates (lower surface temperature) saw significantly lower deposition rates than the higher surface temperatures. CFD simulations using the newly vetted OSU Molten Model were not able to account for the strong decrease in deposition with boundary layer cooling. Therefore, it was posited that the model should eventually incorporate both particle and surface temperatures in its formulation.

Research presented in this thesis focuses on the development of a novel, open-source, automated, and fully parametric non-axisymmetric turbomachinery aero-design-optimization system, NAX. This research forms a part of a collaboration between NASA Glenn Research Center and University of Cincinnati for the development of a Boundary Layer Ingestion based Turbo-electric propulsion system under the NASA N+3 initiative.For the present design, a 1.5 stage aft-mounted Tail Cone Thruster (TCT) unit is considered to ingest the boundary layer flow from the air-frame to increase the overall propulsion efficiency of the aircraft. This results in circumferential distortion (r, &#952)in total pressure, swirl, and meridional flow (PT, &#945,&#981) at the TCT inlet. at the TCT inlet. These non-uniformities causes partial off-design operation of the TCT and a significant departure in its component efficiency and aero-mechanics integrity.A key feature of the NAX design system is its capability to allow for the design/optimize of spanwise (r) and/or circumferentially (&#952) non-axisymmetric 3D blade shapes. It uses a harmonics-based design space parametrization, which offers an extended control in r, &#952 on blade parameters like blade angles, sweep, lean, chord, and thickness distribution. These non-uniformities (&#931 ai sin(&#952 + &#981i)) can be controlled by modifying the magnitude and phase (i) at arbitrary span. NAX is demonstrated with the Inlet Guide Vanes (IGV) for TCT propulsor. This non-axisymmetric IGV is optimized to reduce the downstream rotor incidence and increase IGV performance under 2D distortion. A novel approach of transforming relative flow angle ( &#946) from IGV to rotor frame of reference is used to include the uncoupled rotor effects in the optimization of IGV.The typical multi-fidelity design framework is used to develop TCT using a PT radial distortion at throughflow level and then further optimized for 2D distortion at 3D levels. For design process verification, a manual design is first created using only first mode (a1, &#9811) of circumferential variation of IGV trailing edge angle as input (15&#37 span). A least-squares method (NL2SOL) was then employed using the first mode (a1, &#9811) at 5 spans. Finally, Genetic Algorithm (GA) was used with 2 modes (a1,2,&#9811,2) at span 0, 15 &#37 and 1 mode at the other 3 spans to optimize IGV and reduce stage losses (&#950). Using NAX design system an overall reduction in harmonic content was achieved, 48&#37 using GA (hub), 41&#37 using GA (span 20&#37) and 57&#37 using NL2SOL (span 45 &#37). Swirl, static pressure propagation, and work input at the rotor inlet are also explained, validating the choice of ( &#946) for optimization. Finally, a Finally, a non-axiIGV- axi rotor configuration is optimized for performance at near-stall (&#8773 1&#37 increase in &#951) and design-point using Multi-objective GA (MOGA).This demonstrates the efficacy of NAX-Optimization system for multi-row turbomachinery design.The NAX development and design/optimization framework have enabled a new dimensionality (&#952) into the definition of 6 blade design parameters in harmonics based turbomachinery design space. NAX capabilities can be used for non-axisymmetric end-wall designs, OGV/pylons, and non-axisymmetric volute designs.

Integrally bladed rotors (IBRs), also known as blisks, are important turbomachinery components at the core of gas turbine engines. In operating conditions, IBRs must withstand large centrifugal forces, high temperatures, and blade vibrations caused by upstream flow perturbations. The combination of the cyclic symmetric design of blisks and their inherent mistuning, i.e. small deviations from the nominal properties and geometry, makes them prone to vibration localization, that can lead to vibration amplification of some of the blades in the system. Unlike traditional bladed disk assemblies, in which blades are connected to the disk portion using frictional joints, blisks are manufactured in one piece. The lack of frictional interfaces in blisks results in lower damping, which in turn causes even larger vibration amplitudes. The concurrent presence of high steady stresses and vibration, combined with the large numbers of cycles that blisks blades must withstand, can lead to blade failure due to high cycle fatigue. As a result, the study and reduction of blisk vibrations is of paramount importance to ensure the safety and reliability of such components. The study of blisk vibration is currently carried out using reduced order models (ROMs), due to the prohibitive computational cost associated with the repeated solution of large size finite element models. Vibration reduction is commonly pursued using ring dampers, which typically present low effectiveness. A two-fold approach is used in this work to tackle the complex problem of vibration characterization and mitigation in IBRs. First, a set of efficient ROM tools is introduced to be used in the design phase. Then, novel damping mechanisms are proposed for vibration reduction in operating conditions. A conditioning technique for projection-based ROMs is presented together with examples of applications. Due to the need to model the effects of repairs (blends) or geometric changes during design iterations, models often require the use of distinct meshes for the same component. A novel method is thus developed to overcome computational issues caused by the use of multiple meshes in ROMs for mistuned blisks. Both methods preserve the accuracy of the ROM and at the same time reduce the overall computational cost. In addition to this, a method is presented to efficiently capture the effects of prestress in geometrically mistuned blisks. Different cases and blend types are examined to prove the generality of the approach, showing excellent agreement with the full order model. The last three chapters are entirely dedicated to novel ring dampers concepts, in which the core idea behind tuned vibration absorbers is leveraged to substantially increase damper effectiveness. This allows for higher levels of energy dissipation by frictional nonlinearities, making it possible for the damper to reduce vibrations even when placed in a region with little blisk-alone motion. First, a preliminary computational study is introduced to prove the validity of the concept. Then, a test rig is presented to experimentally demonstrate the effectiveness of the concept. Finally, a novel damper design is presented and studied using state-of-the-art computational tools.

In addition to their efficiency, the versatility of gas turbines is a major driver in their development for aero engines and power production. Advanced materials, cooling and thermal management are the main factors driving the reliable operation of gas turbines with main gas temperatures above their metal melting temperature. While adequate cooling is highly necessary, it comes at a cost. An improved understanding of heat transfer can reduce cooling requirements for components and shift the trade-off between increasing turbine inlet temperatures and associated cooling requirements for higher efficiencies. Improvements to design methods and new manufacturing techniques-such as neural network-based computational fluid dynamic optimisations and 3D printing-broaden the design space for engineers and facilitate increasingly integrated component designs. This results in a strong need for cost- and time-effective experimental testing and verification of numerical predictions. Due to the complexity of measuring high-resolution heat transfer on rotating components in gas turbine research facilities, relatively little experimental data are available in the open literature. Instead, most investigations are based on experiments in linear cascades with low geometrical complexity and simple inflow conditions. In particular, the unsteady flow field and interactions between main flow and cavity flows is not representative in such experiments. Established techniques used on rotating components such as thin-film gauges provide only single-point measurement data. This thesis contributes to the experimental setup and instrumentation for high-resolution heat transfer measurements on rotating turbomachinery components by improving a novel setup using a high-speed infrared-based technique. The rotor hub endwall heat transfer coefficient and purge flow cooling effectiveness of an endwall with contouring extended into the disc cavity is investigated for various purge flow injection rates. Using numerical simulations, the capabilities and challenges on heat transfer are assessed. The work in this thesis was performed within a joint industrial and academic research project to investigate the effect of advanced airfoil designs on the aerothermodynamics of gas turbines with cooling flows, with a special focus on heat transfer. Experimental measurements were performed in a state of the art and highly accurate axial turbine research facility. Aerodynamic measurements using pneumatic five-hole probes and fast response aerodynamic probes were performed to characterise the flow field, define boundary conditions and validate the numerical simulations. Moreover, a previously developed heat transfer measurement setup was improved for highly curved surfaces and rotor blade tip measurements. Insert-based custom-made thin-film resistive heaters based on chemical deposited nickel on a low thermal conductive substrate were manufactured and integrated on an aluminium bladed disk rotor to create controlled heat flux boundary conditions. The pressure and temperature on the rotor and the power for the heating platforms were acquired and controlled using a rotating data acquisition system mounted on the rotor. The test facility was modified to integrate the injection of dry cold air from a secondary loop to create a temperature difference between the injection cooling air and the main flow. A modular tip instrumentation setup was introduced to allow tip cooling integration for bladed disk rotors and simultaneous tip heat transfer measurements in a rainbow rotor approach with multiple geometries in a single test run. Complementary numerical investigations were performed using a Navier-Stokes solver for turbomachinery applications on a high-performance computing cluster. The developed technique for nickel thin-film resistance heaters was demonstrated to be reliable and versatile for complex geometries with high uniformity in the heat flux produced. The variable blade tip instrumentation was successfully implemented, used for measurement and proven to be robust with over 120 turbine start-ups and nearly 1000 turbine operating hours. For the investigated endwall geometry, local variations in heat transfer coefficient above ±20 % were observed between the blade suction side and the endwall hill for purge flow rate variations between 0:0 % and 1:2 %. The steep contouring hill deflects the main flow into the cavity and promotes a jet-like purge flow ejection from the cavity, thereby limiting the cooled platform area from the purge flow. The effect of purge flow on the endwall is limited to the front part of the platform upstream from the cross-passage migration of the secondary flow structures. Notably, both steady-state and unsteady simulations predicted the distribution of local heat transfer coefficients reasonably well in comparison to the experimental results. Regarding cooling effectiveness, an overestimation by the steady-state solution was observed due to suppression of the unsteady interaction between the main flow and cavity as well as the mixing of the two. To the best of the author’s knowledge, this study provides the first high-resolution heat transfer measurements at rotor blade tips performed in a rotating facility using a high-speed infrared camera and heat flux controlled boundary conditions. A unique data set of heat transfer coefficient data for various rotor blade tip geometries-with and without film cooling-has been acquired. Pronounced geometrical features on the rotor platform, such as upstream and cavity extended endwall contouring, can introduce regions of increased heat transfer that are not caused by secondary flows structures created in the passage and should be considered in the design phase. Due to the unsteady nature of the cavity and main flow interaction, unsteady simulations are required to correctly predict the cooling effect of purge flow injection on the hub endwall.

Blending is a method of fan and compressor blade repair. The goal of the blending process is to remove stress concentration points such as cracks and nicks along the leading, trailing, or tip edges of the blade. The stressed areas are typically removed by grinding or cropping away the surrounding material. For integrally bladed rotor (IBR) disks, repairing a damaged blade is much more economical than replacing the entire disk. However, the change in shape of the blade will change the local aerodynamics and result in mistuning, both structurally and aerodynamically. In a worst case scenario, the change in the aerodynamic forces acting on the blades could lead to either flutter or resonant fatigue failure. This study focused on the effects of the unsteady aerodynamic loading of blended blades. To examine the phenomena in detail a commercial computational fluid dynamics (CFD) code was used to predict the loading of a first-stage, transonic turbofan subjected to varying degrees of blending. The study revealed the effects of the blend are not limited to the blended blade, as changes in steady and unsteady pressure loading was predicted on the other nonblended blades on the rotor. A maximum steady sectional loading change of 7% at 99% span on the blended blade for the maximum sized blend case was predicted. The unsteady analysis found an 18% maximum change in unsteady 1st mode sectional loading on the blade adjacent to the blended blade for the 18 engine order excitation at 90% span. Therefore, the potential for a serious aeromechanics issue exists which requires an unsteady aerodynamic analysis to be performed. It is difficult to determine the true aeromechanical effects without a coupled aerodynamics/structural analysis, but this aerodynamic loading investigation suggests larger blends could be safely used and should be analyzed.

The efficiency of gas turbines is known to increase with the exit temperature of the combustor gases. However, this temperature is limited by the melting point of various equipment downstream of the combustor. To increase this limit, coolants injected at different locations form low-temperature films on the surfaces of these regions to avoid melting and thermal stress damage. This injection significantly changes the flow field in the vane passage. Therefore, there is a need to study the aerodynamic and thermal effects of combustor, transition duct and passage coolant injection, which can assist the designers of gas turbines to employ better cooling schemes with minimal use of coolant, thus increasing the efficiency and durability of turbines. The study presented in this thesis discusses experimental tests performed to understand the coolant effectiveness for cooling the endwall and vane surfaces of a nozzle guide vane cascade. The test section contains an engine representative combustor-turbine interface along with a contoured endwall. High turbulence intensity as well as high Reynolds numbers are achieved in the facility to closely simulate engine conditions. In addition to recording the surface effectiveness values, in-field thermal and aerodynamic measurements were taken. It has been previously discovered that the effusion and louver coolants, used to cool the combustor section, can also be credited in cooling the endwall and the vane surfaces. Therefore, this study is helpful to understand the effects of changing the mass flow ratios of different coolants injected upstream of the passage. Especially, louver coolant injection in the combustor transition duct region, due to its location and injection angle, is suspected to provide significant passage endwall cooling. In-field measurements provide insight to the coolant transport through the passage and its interaction with the mainstream. As the net combustor coolant momentum is higher than the film coolant momentum, the changes in the flow field due to its injection are more significant and are sustained to the end of the passage, giving more streamwise coverage than with a conventional film cooling setup. The film coolant mass flow ratio is varied in this study to see how the interaction of different coolants changes with changes of their injected momentum. The measurements reveal that the upstream coolant flow helps in keeping the film coolant attached to the endwall for most of the passage and also keeping the film thickness fairly uniform in the pitchwise direction. An increase in louver coolant mass flow rate shows higher endwall cooling effectiveness. The velocity contours show that a dominant vortex is present due to combustor coolant injection and that the passage vortex was, although present, diminished due to this other vortex. While changing the louver coolant mass flow rate did not change the intensity of these vortices, changes in the film coolant mass flow rate increased the intensity of the dominant vortex near the pressure surface. Also, the presence of this new vortex helps in film cooling the pressure surface, a region where, conventionally, the least amount of coolant coverage is recorded.

A primary source of periodic unsteadiness in low-pressure turbines is the wakes shed from upstream blade rows due to the relative motion between adjacent stators and rotors. These periodic perturbations can affect boundary layer transition, secondary flow, and loss generation. In particular, for high-lift front-loaded blades, the secondary flowfield is characterized by strong three-dimensional vortical structures. It is important to understand how these flow features respond to periodic disturbances. A novel approach was taken to generate periodic unsteadiness which captures some of the physics of turbomachinery wakes. Using stationary pneumatic devices, pulsed jets were used to generate disturbances characterized by velocity deficit, elevated turbulence, and spanwise vorticity. Prior to application in a turbine flow environment, the concept was explored in a small developmental wind tunnel using a single device. The disturbance flowfield for different input settings was measured using hot-film anemometry and Particle Image Velocimetry. Insight was also garnered on how to improve later design iterations. With an array of devices installed upstream of a linear cascade of high-lift front-loaded turbine blades, settings were found which produced similar disturbances at varying frequencies that periodically impinged upon the leading-edge region. These settings were used to conduct an in-passage secondary flow study using high-speed Stereoscopic Particle Image Velocimetry. Results demonstrated the application of the periodic unsteadiness generator but found minor changes to the passage vortex. The vortex rotational strength decreased, and migration increased with increased perturbation frequency. Fourier analyses found the PV to be responsive at the actuation frequency with phase-locked ensemble-averaged data revealing that the disturbance periodically caused the PV to lose rotational strength. However, at the tested discrete frequencies, the vortex did not become locked-in and was defined by an erratic time-character similar to without the disturbances.

In turbomachinery design, complex internal flows give rise to significant losses and blockage whose effects are difficult to properly analyze without detailed computational fluid dynamics (CFD) methods or experiments. In a typical design method, CFD is used in conjunction with simpler throughflow or cascade codes to hasten the process. However, the lesser physical accuracy of the design codes demands the inclusion of models to improve the accuracy of the throughflow codes. This thesis aims to use CFD data to generate improved loss and blockage models for a 2D compressor throughflow code by matching throughflow data to CFD data using optimizations. This analysis gives insight into the connection between 3D flow features and design blockages and losses. This in turn forms the foundation for the behaviors required in the models for shocks, separations, and tip clearance flows, the implementation of which will improve the physical accuracy of the design code.

The low-pressure turbine is an important component of a gas turbine engine, powering the low-pressure spool which provides the bulk of the thrust in medium- and high-bypass engines. It is also a significant fraction of the engine weight and complexity as it can comprise up to a third of the total engine weight. One way to drastically reduce the weight of the low-pressure turbine is to utilize high lift blades. To advance high-lift technology, the Air Force Research Laboratory (AFRL) designed the L2F blade profile, which was implemented in the linear cascade at AFRL/RQT’s low speed wind tunnel facility. The L2F blade has very high lift and an excellent midspan performance, however, it was previously demonstrated to generate significant losses in the endwall region. These losses are primarily driven by the complex time-dependent three-dimensional vortical structures present in the region of the junction of the blade and the endwall, dominated by the Passage Vortex (PV). Aerodynamic flow control is one way to mitigate these losses. Previously, a pulsed endwall blowing system was implemented in the endwall region of the L2F blade which produced a loss reduction. This loss reduction was dependent on the pulsing frequency. In this research, the vortical structures for the baseline flow were characterized with respect to time. The time dependent behavior of the passage vortex motion, location, and strength were found for each pulsing frequency to determine a relationship with total pressure loss reduction. The flow through the passage of the tunnel was characterized with respect to time using high-speed stereoscopic particle image velocimetry. The flow for each test condition was characterized using Q-criterion to determine the strength of the passage vortex and its time dependent behavior. It was found that the passage vortex loses and gains strength in an unsteady manner at time scales between 1.9 < ΔT+ < 6.7. The largest total pressure loss reduction was found to correspond to the pulsing frequency at which the passage vortex was the weakest and moved the closest to the suction surface of the passage

The primary scope of this thesis is to contribute to improving the aerothermal performance of turbine sections in modern gas turbines for electricity production and aircraft propulsion in aero engines. The high-pressure turbine components in such machines are typically subjected to significant thermal loads induced by the continuously increasing turbine entry temperatures. Sensitive components, such as rotor disks and blades, are consequently supplied with cooling flows that, in turn, tend to have a detrimental impact on the aerodynamic performance of the turbine. In that context, geometrical features, like the rim seals and the rotor blade tips, bear great potential for optimization. Furthermore, the designers aim to reduce the number of turbine blades to cut the cooling requirements. Within that design space, this work targets providing guidelines for optimizing the rim seal space and rotor blade tips to efficiently use and reduce the expensive cooling flows while maintaining high aerodynamic performance levels. Potential unfavorable concomitant effects on the turbine characteristics (such as turbine noise) are elucidated in the context of the optimization process. The present work was carried out within the scope of a joint academic and industrial research effort to improve the aerothermal performance of high-pressure turbines by integrating novel geometrical features such as advanced rim seal designs and rotor blade tips in combination with non-axisymmetric end wall contouring. In total, three different 1.5-stage turbine configurations with multiple sub-assemblies, each consisting of specific rim seal and tip designs, were probed for different cooling flow rates in the axial turbine facility “LISA” at ETH Zurich. A unique experimental dataset is presented consisting of inter-row pneumatic and fast-response probes as well as cavity wall-mounted pressure and temperature measurements. The unsteady flow perturbations inside the delicate rim seal space were resolved using miniature-sized, piezo-resistive pressure transducers installed both on the stator- and rotor-sided hub cavity walls. A purpose-made hub cavity heat transfer setup consisting of thin-film heater and double-sided heat flux gauges allowed for changes in the thermal boundary conditions during turbine operation and thereby offered the determination of heat transfer and ingestion quantities. A modular multipurpose rotor onboard telemetry system was used, enabling the acquisition and subsequent transfer of pressure and temperature data from the rotor relative frame of reference. The test rig capabilities were extended by a low-temperature bypass flow system, which facilitated increasing the temperature difference between cavity fluid and main annulus flow. A modular, bladed-disk design was introduced to allow for a rainbow rotor setup featuring different blade tip designs for the same test run. The experimental work was complemented by extensive unsteady computational fluid dynamics simulations using the in-house developed explicit solver MULTI3. The complex flow field in the rim seal space was traced by fast-response instrumentation and unsteady calculations. The significance of specific non-synchronous low-frequency flow perturbations (cavity modes) for the aerothermal turbine design was elaborated, indicating that the main annulus ingestion behavior, as well as the turbine noise emissions, are directly affected by these flow structures. Transient full-annular numerical modeling presented a reasonable prediction of the low-frequency modes. The extensive probing of different turbine operating conditions and rim seal designs provides measures to attenuate the hub cavity asynchronous flow excitation. A novel rim seal design concept, termed as “purge control features,” is proposed, which was found to improve the aerodynamic performance for an already optimized turbine stage in the presence of rim seal purge flow. In comparison to a baseline case, the absolute stage efficiency increase were experimentally determined to be 0.4% points. Complementary beneficial effects on the ingestion behavior and the attenuation of low-frequency cavity modes were experimentally and computationally quantified. A unique experimental dataset for a combined study of turbine ingestion and hub cavity convective heat transfer is provided for a systematic variation of rim seal purge flow rates. The sensitivity analysis of the ingestion and heat transfer with respect to purge air highlights the importance of cooling flow rates to the overall improvement of the aerothermal performance of a turbine stage. An increase in the local convective heat transfer coefficient by a factor of 2–3 on the rotor-sided hub cavity wall was detected for a 1% increase in purge flow injection rate. The resultant heat transfer coefficients were compared to existing empirical correlations, indicating the limitations of such approaches in providing accurate predictions. Extensive probing of a substantially reduced blade count rotor featuring an optimized blade and end wall design revealed the potential to improve the turbine stage performance by 0.4% points. However, the achievements were compromised by a significant increase in turbine tonal noise emissions by up to 13dB. An additional absolute performance gain, up to 0.1% points, was seen by integrating a set of advanced blade tip geometries that were experimentally cross-compared using a rainbow rotor setup. The combined absolute stage performance gain was found to be 0.5% points for a 22% reduction in rotor blade count, resulting in an overall higher aerothermal performance.

Steam turbines hold the largest share of electricity production worldwide. In order to meet growing energy demands, steam turbine designers strive for increased power output, improved efficiency, and longer operational lifespan. Modern steam turbine designs employ large exit annulus area with long blades of up to 60 inches in the last stage of a low-pressure steam turbine in order to achieve these objectives. Increase in last stage blade height introduces supersonic flows in the tip span of the last stage rotor inlet. Such flows are subject to high unsteadiness induced by shock waves. In addition, condensation, droplet formation and unsteady blade loads in the rotor transonic tip region introduce a very complicated flow for numerical computations. Therefore, optimization of blade stacking in the last stage of a transonic low-pressure steam turbine is one of the most delicate and time-consuming parts of the design process. This requires design modifications including blade sweep, lean or twist. The experiments for scaled geometries are nevertheless very expensive and designers have to rely on time accurate computational fluid dynamics. The accuracy of computations is extremely critical in order to guide optimization algorithms and designers to the most viable design. The time-accurate computations are an order of magnitude more expensive than steady state. During design optimization, detailed geometrical features are excluded in order to achieve realistic computational runtime at the cost of accuracy. The geometrical features mostly excluded are hub-tip cavities, seals, part span snubbers, full span shroud, blade count modification or exclusion of upstream or downstream stages. Full-scale multi-stage model with all-inclusive geometrical features results in very large meshes of up to one billion mesh nodes. The proposed meshes must have matching block interfaces throughout the mesh in order to keep a second-order accuracy in space and time posing additional requirements of a very fine mesh in order to resolve high blade twist, shroud connections, and cavities-seals in the flow path. In the case of low-pressure steam turbines, the steam transition from superheated to condense in penultimate and last stages. This necessitates the inclusion of steam modeling at the least for the prediction of wetness as well as numerical accuracy. This results in additional computational resource requirements posing a very challenging computing problem. The recent widespread use of modern general-purpose graphics processing units (GPUs) for scientific computing provides a possibility to scale time-accurate computational fluid dynamic solvers for such challenging problems. A steady decline in graphics processing unit costs and at the same time improvements in throughput and onboard memory gradually allow hybrid high performance computing cluster as a viable option for the design engineering process. The key objective of this work is to improve the aerodynamic efficiency of a modern low-pressure steam turbine, using carefully tailored stator stacking, numerically on a cluster of GPUs and explain underlying time-resolved flow mechanisms. For this purpose, in-house Unsteady Reynolds Averaged Navier-Stokes (URANS) solver MULTI3 is developed for multi-GPU supercomputing clusters with additional equilibrium steam modeling, able to handle full-scale model. In this work, the typical scale of about 386 million mesh nodes for a full annular four-stage model running on 114 Pascal P100 GPUs with time-resolved equilibrium steam modeling is reported with a convergence run-time of less than a week. For design optimization, a typical scale of 57 million mesh nodes including tip-cavities and seals for the last two stages is achieved within a week of run-time on 17 GPUs. Last two-stage computational model of 16.9 million mesh nodes excluding tip cavities converge within 21 hours of reported runtime on two GPUs. Further, a key understanding is developed on computational accuracy versus cost with each geometrical or modeling approximation and guidelines are proposed for a suitable compromise of numerical approximation in a low-pressure steam turbine design optimization process. The results show the transonic tip region as the main source of entropy loss in the last stage of the transonic low-pressure steam turbine. The proposed stator stacking with an increase in tip axial gap (forward curved sweep) and throat-to-pitch ratio variation designs show aerodynamic total-to-total efficiency improvements of up to 1.3% and 1.1% respectively. The key improvements are seen in supersonic flow expansion in the last stage rotor tip driven by a reduction of relative inlet Mach. The flow unsteadiness in the transonic stator-rotor tip gap is primarily driven by leading edge bow shock, however as found in this work, also depends on the axial gap. Despite an increase in unsteadiness by the closing throat (stator twist design) in the last stage stator tip, a decrease in relative Mach improves efficiency by improvements in the flow expansion in the supersonic rotor tip airfoils. The control of throat-to-pitch ratio is found to be the most effective way to control reaction variation and mass flow redistribution in both the stator and rotor. The analysis shows that at least penultimate stage is compulsory in the computational model along with tip cavities and seals for the last stage optimization as any change in throat area as a result of stacking modification introduce a change in stage pressure ratio, workload, and reaction for both stages.

There is a continued need to further capabilities of turbomachinery flow design and analysis method tools, both in terms of the flow modeling fidelity (e.g. moving towards DES/LES) and temporal and spatial accuracy of the numerical discretization methods. To enhance solution accuracy with high-order methods, an implicit solver using the Discontinuous Galerkin (DG) discretization on unstructured meshes has been developed. The DG scheme is favored chiefly due to its distinctive feature that a higher-order accuracy being achieved through simple internal sub-divisions of a given mesh cell. It thus can considerably relieve the burden of mesh generation for local refinement of complex configurations. The implicit formulation is well suited for an unstructured mesh solver, for which some well-established solution techniques for explicit methods (e.g. multigrid methods) are less effective. The RANS equations in conjunction with the Sparlart-Allmaras one equation turbulence model are discretized using a high-order DG method on unstructured meshes. The all-speed AUSM+-up scheme is applied for the inviscid flux calculations and the BR2 scheme is implemented for the viscous flux discretization. An implicit time integration is adopted and the system linear equations are solved at each time step by using a preconditioned GMRES iterative algorithm with the ILU0. A modification of the turbulence model is also adopted to enhance the robustness for high-order turbulent flow simulations. The validity and effectiveness of the present developed method have been examined in several test cases, for both validating the method implementation and illustrating salient features and characteristics of the methodology. The flows around a cylinder solved up to 5th-order of accuracy on an extremely coarse mesh illustrate the need for high-order geometrical representation in relation to the high-order flow field discretization. A laminar boundary layer and the vortex shedding after a cylinder validate both the steady and unsteady modelling capabilities with the implicit time integration method. A calculated turbulent boundary layer on flat plate demonstrates the capability of a high-order DG in capturing the laminar sub-layer in a coarse mesh (y+>20) without a wall function, in a clear contrast to a conventional 2nd-order scheme. The all-speed flows past an airfoil show the distinct capability of AUSM+-up scheme on solving a wide range of flow regimes. A main focal test case of the present work is a high pressure turbine cascade, where the flow losses are strongly influenced by the boundary layer transition on the blade surface and the secondary flow development in the endwall region. In this case the RANS solutions consistently over-predict the losses, prompting further modelling considerations. A higher-order solution does indicate higher resolution for more detailed 3D secondary flow structures. It is however more remarkable that both overall and distributed losses of the unsteady laminar solutions on a coarse mesh are consistently better than the RANS solutions. The results indicate that for the turbine blade flows similar to the one presently tested, a direct unsteady laminar solution on a coarse mesh (far coarser than those typically required for LES) may be further explored.

The predictable effect of surface curvature on the contiguous streamtube allows for the use of geometric curvature as a direct and aerodynamically meaningful parametric model to generate airfoil geometry. A novel and parsimonious parameterization technique driven by specifications of normalized meanline second derivatives, which is related to curvature, and a superimposed thickness distribution which explicitly eliminates or minimizes unintentional oscillations in curvature, is resented. The focus on smooth curvature control is inherently relevant to both internal and external flow geometries, enabling a unified method for generating both isolated airfoils and cascade sections. This technique is implemented and included in T-Blade3 which is an existing in-house open-source code. The underlying methodology for construction of camber-line is entirely analytical ensuring speed of execution. Two different thickness distributions are presented, one based on specifications of thickness B-spline control points, and another based on specifications of exact thickness. The B-spline thickness method requires a simple implementation and executes faster, but cannot impose an exact value of thickness at a specified location. The exact thickness method is capable of imposing specified thickness values both chord-wise and span-wise, while quantifying and optimizing the quality of the thickness curve based on curvature. Consequently, unintentional bumps on the airfoil and oscillations in curvature are eliminated or minimized. The parameterization ensures curvature and slope of curvature continuity on the airfoil surface which are critical for smooth surface pressure distributions. Consequently, losses due to unintentional pressure spikes are minimized and likelihood of separation reduced, resulting in a class of high-performance airfoils. The direct relationship between the parameterization and surface aerodynamics is demonstrated for both isolated and cascade airfoils. A framework for 2D airfoil shape optimization is developed which uses a parallelized Genetic Algorithm implementation from the US Department of Energy optimization system, DAKOTA. As a demonstration of the parameterization capability, optimization is performed to maximize lift-to-drag ratios of an isolated airfoil over a range of angles of attack at two Reynolds numbers using Genetic Algorithm. The tool-chain uses XFOIL to evaluate drag polars. Two different cases are presented, each using a different thickness parameterization. The maximum thickness to chord ratio is fixed at 21%, enabling direct comparison of the optimized shapes with the S809 airfoil, which is used as a comparison benchmark. Drag improvements ranging from 17% to 55% for Cl between 0.3 and 1, for Reynolds numbers of 7*10^5 and 9.5*10^5 were achieved by the optimized airfoil using the B-spline based thickness parameterization, while improvements ranging from 11% to 52% were achieved by the optimized airfoil using exact thickness parameterization for the same range of Reynolds numbers and Cl .

Sweeping jet impingement cooling was investigated in a gas turbine nozzle guide vane design with an engine-relevant Biot number of 0.3. Sweeping jets were created with fluidic oscillators and were compared to steady jets produced by cylindrical orifices (with length-to-diameter ratio of 1), the current state-of-the-art in engine designs. Experiments were performed in a low speed linear cascade with additively manufactured test pieces. The impingement cooling geometries were examined at multiple coolant mass flow rates and freestream turbulence intensities. The overall effectiveness of each cooling geometry was calculated using thermocouple measurements of the freestream and coolant temperatures, and infrared thermography measurements of the vane external surface temperature. A computational thermal inertia technique was used to determine the internal Nusselt numbers. The heat transfer provided by steady impinging jets produced a higher overall effectiveness and Nusselt number in the leading edge geometry. The sweeping jets provided more uniform heat transfer, reducing thermal gradients near the stagnation point. Pressure drop across each jet geometry was measured at a range of applicable mass flow rates. Fluidic oscillators were shown to create similar pressure drop to circular orifice holes when additive manufacturing abilities were fully incorporated in the nozzle guide vane internal cooling designs.

Significant effort in the current technological development of aircraft is aimed at improving engine efficiency, while reducing fuel burn, emissions, and noise levels. One way to achieve these is to better integrate airframe and propulsion system. Tighter integration, however, may also cause adverse effects to the flow entering the engines, such as total pressure, total temperature, and swirl distortions. Swirl distortions are angular non-uniformities in the flow that may alter the functioning of specific components of the turbomachinery systems. To investigate the physics involved in the ingestion of swirl, pre-determined swirl distortion profiles were generated through the StreamVane method in a low-speed wind tunnel and in a full-scale turbofan research engine. Stereoscopic particle image velocimetry (PIV) was used to collect three-component velocity fields at discrete planes downstream of the generation of the distortions with two main objectives in mind: identifying the physics behind the axial development of the distorted flow; and describing the generation of the distortion by the StreamVane and its impact to the flow as a distortion generating device. Analyses of the mean velocity, velocity gradients, and Reynolds stress tensor components in these flows provided significant insight into the driving physics. Comparisons between small-scale and full-scale results showed that swirl distortions are Mach number independent in the subsonic regime. Reynolds number independence was also verified for the studied cases. The mean secondary flow and flow angle profiles demonstrated that the axial development of swirl distortions is highly driven by two-dimensional vortex dynamics, when the flow is isolated from fan effects. As the engine fan is approached, the vortices are axially stretched and stabilized by the acceleration of the flow. The flow is highly turbulent immediately downstream of the StreamVane due to the presence of the device, but that vane-induced turbulence mixes with axial distance, so that the device effects are attenuated for distances greater than a diameter downstream, which is further confirmed by the turbulent length scales of the flow. These results provide valuable insight into the generation and development of swirl distortion for ground-testing environments, and establishes PIV as a robust tool for engine inlet investigations.

Numerical methods for computational physics have been applied for many years in the fields of turbomachinery and acoustics. The computational approach to addressing problems in these fields has strongly influenced improvements in performance and advancements in understanding for the systems and physical processes that govern such applications. Particularly within the turbomachinery community, the spectrum of numerical methods being applied is dominated by second-order accurate finite-volume and finite-difference discretizations of the governing equations. These have been quite successful and their robustness has been important in their adoption within industrial engineering as tools for design and analysis. At the same time, such approaches are very dissipative and require dense computational grids to resolve sharp features and capture wave propagation. As problems become more complex and inter-related, the numerical methodologies for analysis tools that are used to address such problems and inform solutions must improve in their fidelity, accuracy, and mathematical rigor.At the same time, high-order finite-element methods have experienced significant attention in the computational fluid dynamics community for their mathematical formalism, localized approach for obtaining high-order accuracy, and amenability to adaptation of the numerical grid and accuracy of the numerical approximation. The challenges of such approaches are to achieve efficiency and stability for the numerical method. The result of many research efforts in this area has pushed the applicability of high-order finite-element methods into many fields and they are approaching the point where they might soon find routine application to industrial problems for engineering design and analysis. This dissertation details the development and application of an implicit discontinuous Galerkin method to applications in turbomachinery and acoustics. This is carried out for the purpose of advancing the applicability of the discontinuous Galerkin method to problems in these fields. A selection of the developments included in this work include a study on the efficiency of automatic differentiation for the discontinuous Galerkin method, adoption of a new approach for computing distance fields based on solving a p-Poisson equation, the development of a new adiabatic no-slip boundary condition for moving walls, and the development of implicit nonlocal nonreflecting boundary conditions along with a new stabilizing correction that significantly improves their efficiency and robustness. Mixing-plane interfaces for turbomachinery applications were also developed for the first time within a discontinuous Galerkin discretization. Finally, the numerical method that was created in the context of this work was applied to several turbomachinery problems. These problems include steady, Reynolds-Averaged Navier-Stokes (RANS) calculations for a two-dimensional stator, a three-dimensional stator, and a quasi three-dimensional 1.5 stage turbine using nonreflecting mixing-plane interfaces. These contributions and their application significantly advance the state-of-the-art for high-order finite-element methods applied to computational fluid dynamic problems in turbomachinery.

Turbine bladed disks (blisks) which constitute critical components of most modern turbomachinery, are known for their complex vibratory behavior. A good blisk design necessitates the knowledge of the dynamic behavior of the component in different regimes of operation. However, small parametric uncertainties and inter-blade structural variations called mistuning are known to affect blisk dynamics drastically under certain conditions. Introduced by manufacturing tolerances, thermal stresses, wear and other causes, these variations make the design process especially challenging. The dynamics of blisks often also involve nonlinearities, such as those arising from friction damping structures such as shrouds or under-platform dampers or from nonlinear damping coatings adhered to the blades. Since it is not possible to eliminate uncertainties due to mistuning, the designer must ensure safe operation within the field of uncertainty of all the design parameters. Presently, the widely accepted method for obtaining a comprehensive understanding of the effects of these uncertainties is to conduct probabilistic analyses with many simulations of different systems generated by applying random mistuning to the nominal tuned system. Finite element modeling techniques when applied in isolation prove too cumbersome and time-intensive to use for this task. To make these simulations computationally feasible, reduced order models which accurately capture the nonlinear mistuned dynamics while being computationally tractable must be developed. This work presents model reduction methods for blisk dynamics involving two specific nonlinear mechanisms. An adaptive microslip projection method is developed to reduce the dynamics of blisks with frictional contacts characterized by complex stick-slip behavior. A reduced order modeling technique developed for blisks with nonlinear coatings using an amplitude dependent mistuning framework to simulate the dynamics is presented also. Both reduced order models are found to provide significant computational time savings, while retaining high accuracy in comparison to their higher order baseline models. Another practical challenge for accurate nonlinear dynamic simulation of blisks, is the identification of physical parameters which represent the physics of the system and are thus employed by both the full and reduced order models. Contact parameters are identified from experimental data using optimization techniques and are used to validate the reduced order model. The specific effects of variations in these contact parameters on the dynamic response of the blisk are also studied by carrying out probabilistic analyses.

This research presents a system design and analysis, and transient \textit{control} and optimization of solid oxide fuel cell-gas turbine (SOFC-GT) hybrid systems. The main features of SOFC-GT power systems include high efficiency, virtually zero emission of criteria pollutants, and low acoustic signature compared to conventional power production technologies. Focus of this dissertation is the investigation of control strategies and transient performance characteristics of hybrid SOFC-GT systems, using computational fluid dynamics methods.A novel control system is developed for a SOFC-GT system to follow dynamic power demands associated with locomotives while keeping all of the components and system operating variables within acceptable limits of performance. The voltage variation is also greater at lower fuel utilization. The stack temperature difference is highly dynamic at low fuel utilization. In addition, detailed analyses of potential SOFC-GT locomotive production and operation costs in comparison to other low pollutant emitting alternatives is accomplished. SOFC-GT locomotives are projected to produce lower operating costs compared to the catenary-electric alternative, and significantly lower operating costs compared to the battery-electric alternative.A part of this research is devoted to a better understanding of turbulent unsteady flows in gas turbine systems that is necessary to design and control compressors for hybrid fuel cell-gas turbine (FC-GT) systems. In this study, compressor stall/surge analysis for a 4 MW locomotive hybrid solid oxide fuel cell-gas turbine (SOFC-GT) engine is performed based on a 1.7 MW multi-stage air compressor similar to available commercial compressors. Control strategies are designed and evaluated to prevent the operation of hybrid SOFC-GT beyond the stall/surge lines of the compressor. Computational fluid dynamics (CFD) tools are used to provide a better understanding of flow distribution and instabilities near the stall/surge line. Simulation results show the feasibility of using existing industrial compressors in the hybrid SOFC-GT system operation. The results show that a 1.7 MW system compressor like that of Kawasaki gas turbine is an appropriate choice among the industrial compressors to be used in a 4 MW locomotive SOFC-GT with topping cycle design.

Computational fluid dynamics (CFD) simulations can offer a detailed view of the complex flow fields within an axial compressor and greatly aid the design process. However, the desire for quick turnaround times raises the question of how exact the model must be.At design conditions, steady CFD simulating an isolated blade row can accurately predict the performance of a rotor. However, as a compressor is throttled and mass flow rate decreased, axial flow becomes weaker making the capturing of unsteadiness, wakes, or other flow features more important. The unsteadiness of the tip clearance flow and upstream blade wake can have a significant impact on a rotor.At off-design conditions, time-accurate simulations or modeling multiple blade rows can become necessary in order to receive accurate performance predictions. Unsteady and multi- bladerow simulations are computationally expensive, especially when used in conjunction. It is important to understand which features are important to model in order to accurately capture a compressor’s performance.CFD simulations of a transonic axial compressor throttling from the design point to stall are presented. The importance of capturing the unsteadiness of the rotor tip clearance flow versus capturing upstream blade-row interactions is examined through steady and unsteady, single- and multi-bladerow computations. It is shown that there are significant differences at near stall conditions between the different types of simulations.

A 3-D Reynolds-Averaged Navier Stokes (RANS) model of a highly-loaded blade profile has been developed using a commercial CFD code with an unstructured/structured grid and several different turbulence models. The ability of each model to predict total pressure loss performance is examined in terms of the spanwise loss distribution and the integrated total pressure loss coefficient. The flowfield predicted by each model is investigated through comparisons of isosurfaces of Q criterion to previous Implicit Large Eddy Simulation (ILES) results. The 3-equation k-kl-¿ model was shown to provide the most accurate performance predictions for a baseline 3-D LPT geometry, and was then used to analyze the effect a new 3D contoured geometry. The model accurately predicted the qualitative improvement made by the contour by weakening the various vortex structures.

Bladed disks (blisks) used in turbomachinery applications frequently operate under severe forcing conditions, which can lead to high levels of dynamics responses and pre-mature high cycle fatigue (HCF). Small blade-to-blade variations in structural properties, referred to as mistuning, result in strain energy localization, drastically amplifying blisk forced responses, and accelerate HCF. To quantitatively capture the effect of mistuning, it is necessary to develop efficient computational methods to predict the free and forced responses of blisks with various types of mistuning patterns. Due to the high geometric complexity of commercial blisks, full-order finite element (FE) blisk models contain many degrees of freedom (DOFs). Direct structural analyses with such FE models are computationally cumbersome or practically infeasible. Moreover, to prevent blisks from reaching HCF, frictional damping sources are introduced to dissipate vibrational energy and reduce the level of forced responses. Frictional damping is nonlinear in nature, and adds complexity into the blisk systems. Thus, robust and accurate reduced-order models must be developed to predict fast the dynamic responses of blisks with various mistuning and frictional damping sources. The main objective of this study is to develop a framework that involves several novel reduced-order modeling techniques. This framework is capable of efficiently and accurately capturing linear and nonlinear dynamic responses of blisks with small blade material variations, large changes in blisk mass, stiffness, and geometry, and dry friction ring dampers. Moreover, this framework serves as a powerful tool in designing friction dampers with optimal design parameters.

The primary scope of this thesis is to contribute to meeting the challenges associated to the integration of shrouded blades into high speed low pressure turbines of geared turbo-fan engines. Doubling the rotational speed of the low pressure shaft bears considerable advantages in terms of the fuel consumption, the noise emissions, the length as well as the part count compared to conventional multi-spool engine designs. However, the quadratic increase of the centrifugal blade stress with rotational speed has impeded the application of shrouded low pressure turbine blades, which are aerodynamically more favorable and have a higher robustness against flutter instabilities, in current designs due to high cycle fatigue concerns. As a consequence of the reduced error margin in the fatigue design of high speed blade rows, two immediately arising questions for the design of the next gener-ation of geared turbofan engines are addressed in this work: How accurately can current industrial methods predict the unsteady aerodynamic forces in a shrouded low pressure turbine? And: At what performance cost can the mass of a shroud be reduced by plat-form cutbacks to mitigate fatigue issues? The current work was carried out within the scope of a joint academic and industrial re-search effort to elucidate the aerodynamic forcing in a fully shrouded low pressure turbine experimentally and numerically as well as to determine the performance deterioration of different shroud platform geometries. The latest, optimized and tested design of a 1.5-stage fully shrouded low pressure turbine with non-axisymmetric enwalls operating with rim seal purge flow injection represents the baseline for both campaigns. A modular, high resolution and low noise, multi-purpose telemetry system has been developed and inte-grated into the axial research turbine facility “LISA” at ETH Zurich. A unique experi-mental data set comprised of fast response traverse and wall pressure measurements has been acquired. The unsteady surface pressure of both the rotor as well as the second sta-tor has been probed using a total of 60 flush-mounted, ultra-miniature size, piezo-resistive pressure transducers. The aerodynamic forcing functions and the turbine performance have been characterized using a combination of multi-hole pneumatic probe and fast re-sponse aerodynamic probe (FRAP) measurements on the inter-row planes. Leakage flows have been tracked using the unsteady entropy probe (FENT). The comparison of the ex-perimental data with time-accurate RANS simulations highlights the capabilities and lim-itations of commercial solvers and enables the quantification of the prediction error in the aerodynamically induced blade root stress. The unsteady rotor surface pressure in a shrouded, subsonic low pressure turbine is driv-en by convective perturbations and acoustic interaction. Five main forcing functions have been identified: Potential interaction with upstream and downstream rows, convective inflow perturbations and incidence variation in the main flow, secondary flow interaction at the endwalls, inter-blade interaction and leakage and cooling flow interaction. The suc-tion side of the rotor blades is dominated in the front by the relative motion of the first stator’s potential field. The superposition with wake-induced effects results in the peak pressure amplitudes in the order of 13% percent of the rotor pressure drop around the crown at the tip of the blade. The modulation by the potential interaction with the sec-ond stator, whose vane count is changed to differ from the first stator’s for the forcing measurements, results in significant amplitudes at difference and sum frequencies down to the ninth engine order. Contrary to the convective perturbation propagation on the major area of the suction side, the wake-induced surface pressure deficit on the pressure side is reinforced by the acoustic interaction with the adjacent suction side and leads to significantly higher propagation speeds than the flow velocity. Three-dimensional forcing effects have been observed at the hub of the rotor due to the purge air vortex, which is modulated by the first stator potential, and its interaction with the rotor hub passage vortex. Since the phase of local thrust and lift fluctuations are highly sensitive to the purge flow rate, the functionality of the purge flow could potentially be extended from cooling to blade excitation control. Numerical predictions show excellent agreement in vast areas of the turbine flow field. The amplitude and phase characteristic of the dominant rotor surface pressure fluctua-tions are generally captured well. The measurement based correction of the unsteady ro-tor surface pressure enables the quantification of the prediction error of the alternating blade root stresses using fluid-structure-interaction analyses to be 2% to 5%. Turbulence intensity measurements at turbine inlet and transition modelling show potential to fur-ther improve the prediction accuracy. The buildup of non-physical duct modes at turbine exit challenges the standard boundary condition setup in terms of periodicity and fixed exit mass flow. Three different partial tip shroud platform designs have been tested experimentally and compared to a full-shroud baseline. A detailed analysis of the fluid dynamics is carried out and complemented by numerical predictions to provide a starting point for shroud opti-mizations. The leading and trailing edge platforms are cut back separately by 11.6% and 5.3% to isolate the effect of each modification. The final rotor then features a combined partial shroud. The performance reduction for both isolated cutbacks equals 0.7% relative to the baseline, while the detrimental effects add up to 1.1% for the combined cutback. The main loss mechanism for all cutbacks is the intensification of fluid exchange with the shroud cavities and the re-entry into the main flow. This results in increased blockage by the tip passage vortex for the leading edge cutback, while the injection of leakage fluid in-to the suction side boundary layer of the rotor results in a flow underturning for the trail-ing edge cutbacks. The usage of shroud cutbacks proves to be an effective measure to at-tenuate the noise emission and excitation potential of non-synchronous, low frequency cavity modes by up to 15dB at the exit of the turbine.

Future civil transport designs may incorporate engines integrated into the body of the aircraft to take advantage of efficiency increases due to weight and drag reduction. Additional increases in engine efficiency are predicted if the inlets ingest the lower momentum boundary layer flow that develops along the surface of the aircraft. Previous studies have shown, however, that the efficiency benefits of Boundary Layer Ingesting (BLI) inlets are very sensitive to the magnitude of fan and duct losses, and blade structural response to the non-uniform flow field that results from a BLI inlet has not been studied in-depth. This project represents an effort to extend the modeling capabilities of TURBO, an existing rotating turbomachinery unsteady analysis code, to include the ability to solve the external and internal flow fields of a BLI inlet. The TURBO code has been a successful tool in evaluating fan response to flow distortions for traditional engine/inlet integrations. Extending TURBO to simulate the external and inlet flow field upstream of the fan will allow accurate pressure distortions that result from BLI inlet configurations to be computed and used to analyze fan aerodynamics and structural response. To validate the modifications for the BLI inlet flow field, an experimental NASA project to study flush-mounted S-duct inlets with large amounts of boundary layer ingestion was modeled. Results for the flow upstream and in the inlet are presented and compared to experimental data for several high Reynolds number flows to validate the modifications to the solver. Once the inlet modifications were validated, a hypothetical compressor fan was connected to the inlet, matching the inlet operating conditions so that the effect on the distortion could be evaluated. Although the total pressure distortion upstream of the fan was symmetrical for this geometry, the pressure rise generated by the fan blades was not, because of the velocity non-uniformity of the distortion. Total pressure profiles at various axial locations are computed to identify the overall distortion pattern, how the distortion evolves through the blade passages and mixes out downstream of the blades, and where any critical performance concerns might be. Stall cells are identified that are stationary in the absolute frame and are fixed to the inlet distortion. Flow paths around the blades are examined to study the stall mechanism. Rather than a static airfoil stall, it is observed that the non-uniform pressure loading promotes a three-dimensional dynamic stall. The stall occurs at a point of rapid incidence angle oscillation, observed when a blade passes through the distortion, and re-attaches when the blade leaves the distortion.

In many regions of the high-pressure gas turbine, film cooling flows are used to protect the turbine components from the combustor exit hot gases. Endwalls are challenging to cool because of the complex system of secondary flows that disturb surface film coolant coverage. The secondary flow vortices wash the film coolant from the surface into the mainstream significantly decreasing cooling effectiveness. In addition to being effected by secondary flow structures, film cooling flow can also affect these structures by virtue of their momentum exchange. In addition, many studies in the literature have shown that endwall contouring affects the strength of passage secondary flows. Therefore, to develop better endwall cooling schemes, a good understanding of passage aerodynamics and heat transfer as affected by interactions of film cooling flows with secondary flows is required. This experimental and computational study presents results from a linear, stationary, two-passage cascade representing the first stage nozzle guide vane of a high-pressure gas turbine with an axisymmetrically contoured endwall. The sources of film cooling flows are upstream combustor liner coolant and endwall slot film coolant injected immediately upstream of the cascade passage inlet. The operating conditions simulate combustor exit flow features, with a high Reynolds number of 390,000 and approach flow turbulence intensity of 11% with an integral length scale of 21% of the chord length. Measurements are performed with varying slot film cooling mass flow to mainstream flow rate ratios (MFR). Aerodynamic effects are documented with five-hole probe measurements at the exit plane. Heat transfer is documented through recovery temperature measurements with a thermocouple. General secondary flow features are observed. Total pressure loss measurements show that varying the slot film cooling MFR has some effects on passage loss. Velocity vectors and vorticity distributions show a very thin, yet intense, cross-pitch flow on the contoured endwall side. Endwall adiabatic effectiveness values and coolant distribution thermal fields show minimal effects of varying slot film coolant MFR. This suggests the dominant effects of combustor liner coolant. show dominant effects of combustor liner coolant on cooling the endwall. A coolant vorticity correlation presenting the advective mixing of the coolant due to secondary flow vorticity at the exit plane is also discussed.

The performance of a vapor compression system is known to be affected by the ingestion of liquid droplets in the compressor. In these multiphase flows, the liquid and the vapor phase are tightly coupled. Therefore the interphase heat, mass and momentum transfer as well as droplet dynamics including droplet breakup and droplet-wall interactions play a vital role in governing these flows. Only thermodynamic analyses or two-dimensional mean-line calculations are not sufficient to gain an in-depth understanding of the complex multiphase flow field within the compressor. The objective of this research was to extend the current understanding of the operation of a multistage centrifugal compressor under two-phase flow conditions, by performing three-dimensional computational analysis. In this work, two-phase flow of a single constituent (refrigerant R134a) through a two-stage, in-line centrifugal compressor was analyzed using CFD. The CFD model accounted for real gas behavior of the vapor phase. Novel user defined routines were implemented to ensure accurate calculations of interphase heat, mass and momentum transfer terms and to model droplet impact on the compressor surfaces. An erosion model was developed and implemented to locate the erosion "hot spots" and to estimate the amount of material eroded. To understand the effects of increasing liquid carryover, the mass flow rate of the liquid phase was increased from 1% to 5% of the vapor mass flow rate. The influence of droplet size on the compressor performance was assessed by varying the droplet diameter at the inlet from 100 microns to 400 microns. The results of the two-phase flow simulations were compared with the simulation involving only the vapor phase. Liquid carryover altered the flow field within the compressor, and as a result, both impellers were observed to operate at off-design conditions. This effect was more pronounced for the second impeller. The overall effects of liquid carryover were detrimental to the compressor performance. The erosion calculations showed maximum erosion potential on the blade and shroud of the first impeller. The results from this investigation provided new and useful information that can be used to support improved design solutions.

This thesis details a computational investigation of tone noise generated from a booster(low-pressure compressor) in a fan test rig. The computational study consisted of sets of time-linearized Navier-Stokes simulations in the booster region to investigate the blade-wake interactions that act as the primary noise-generating mechanism for the booster blade-passing frequency and harmonics. An acoustic test database existed with data at several operating points for the fan test rig that was used to compare against the predicted noise data from the computational study. It is shown that the computational methodology is able to capture trends in sound power for the 1st and 2nd booster tones along the operating line for the rig. It is also shown that the computational study underpredicts one of the tones at low power and is not able to capture a peak in the data at the Cutback condition. Further investigation of this type is warranted to quantify the source of discrepancies between the computational and experimental data as the reflected transmisison of sound off the fan through the bypass duct was not accounted for in this study.

Descending maple seeds generate lift to slow their fall and remain aloft in a blowing wind; have the wings of these seeds evolved to descend as slowly as possible? A unique energy balance equation, experimental data, and computational fluid dynamics simulations have all been developed to explore this question from a turbomachinery perspective. The computational fluid dynamics in this work is the first to be performed in the relative reference frame.Maple seed performance has been analyzed for the first time based on principles of wind turbine analysis. Application of the Betz Limit and one-dimensional momentum theory allowed for empirical and computational power and thrust coefficients to be computed for maple seeds. It has been determined that the investigated species of maple seeds perform near the Betz limit for power conversion and thrust coefficient. The power coefficient for a maple seed is found to be in the range of 48 - 54% and the thrust coefficient in the range of 66 - 84%. From Betz theory, the stream tube area expansion of the maple seed is necessary for power extraction. Further investigation of computational solutions and mechanical analysis find three key reasons for high maple seed performance. First, the area expansion is driven by maple seed lift generation changing the fluid momentum and requiring area to increase. Second, radial flow along the seed surface is promoted by a sustained leading edge vortex that centrifuges low momentum fluid outward. Finally, the area expansion is also driven by the spanwise area variation of the maple seed imparting a radial force on the flow.These mechanisms result in a highly effective device for the purpose of seed dispersal. However, the maple seed also provides insight into fundamental questions about how turbines can most effectively change the momentum of moving fluids in order to extract useful power or dissipate kinetic energy.

This work analyzed the performance of two foot pad designs for the compliant finger seal. The double wedge and Rayleigh step pad designs were tested numerically and experimentally. The numerical models were solved using ANSYS CFX and compared to the experimental results. A finite difference based two dimensional model of the fluid film under the pad solving the Reynolds equation coupled with the energy equation was also formulated.The double wedge and Rayleigh step seal pad designs were analyzed using ANSYS CFX. Models matching the design drawings for the actual seal, and models with increased gap heights were created. The models consisted of one or two non-padded feet, padded feet, and the interstices between subsequent feet. Cases were run at rotational speeds of 2 krpm to 12 krpm in 2 krpm increments and at pressures of 5, 20, 25, and 30 psig, which matches the experimental cases. The small gap geometry was run as laminar compressible while the large gap geometry was run as turbulent compressible, and all were run as isothermal. The mass leakage was found to increase with increasing pressure for all models tested. For the double wedge seal and the Rayleigh step seal with the larger gap, the leakage did not change significantly with increasing rotation, but the leakage is affected by the rotation in the Rayleigh step seal with the original gap. Previously unseen compressibility effects, such as potential choking of the flow, were witnessed in the numerical models for the geometries with the larger gaps.Both the double wedge and Rayleigh step seal pad designs were experimentally tested, and the experimental cases matched the parameters of the numerical cases. The results of the experiments on the double wedge seal indicated that as the high side pressure increased, the mass leakage remained relatively unchanged, and the mass leakage was found to decrease with increasing speed. The leakage of the Rayleigh step seal followed a similar trend as the double wedge seal as the high side pressure was increased, however, the leakage of the Rayleigh step seal increased slightly due to higher rotation. The temperature of the seal pads did not change significantly during testing for either seal pad design indicating there was no contact between the rotor and the seal.The results of the numerical models and the experiment were compared. The mass leakage values of the experimental results were an order of magnitude higher than those of the numerical models with the small gap, and compared much better with the large gap model. The higher mass leakage results are likely due to the higher leakage due to the actual seal being warped, and the actual seal possibly having a larger gap height than the model. These issues make comparing the experimental and numeric results difficult.A two dimensional Reynolds equation based solver was written to model the fluid film under a single foot pad. The Reynolds equation is coupled with the energy equation and solved using finite differences. The model was validated by comparing the results with those from a three dimensional Navier-Stokes based model run in ANSYS CFX. The Reynolds based model compared well with the Navier-Stokes model, both qualitatively and quantitatively. Thus, the Reynolds based model provides a fast, effective method of comparing foot pad designs.

There is limited information in the literature on the behavior of shaped film cooling holes fed by crossflow and even less information on the effect of crossflow parameters on film cooling performance. Here, two scaled film cooling models were used to independently vary the crossflow Reynolds numbers in the range of 36,000 to 57,000 and the crossflow velocity ratio from 0.36 to 0.64. Careful attention was paid to controlling physical parameters between comparisons to isolate the effects of internal velocity ratio or Reynolds number on the performance of shaped holes. In the process of controlling the physical parameters of the system, a novel correction for coolant to mainstream density ratio was proposed. The results of this study showed that channel velocity ratio had a larger effect on the film cooling performance of shaped holes than channel Reynolds number. When the mass flux of fluid through the film cooling holes was at the highest and lowest value, increasing the channel velocity ratio decreased the film cooling effectiveness. At a middle mass flux, the outcome was opposite such that an increase in channel velocity ratio resulted in increased effectiveness.

Jet engine eciency goals are driving compressors to higher pressure ratios and engines to higher bypass ratios, each one driving to smaller cores. This is leading to larger tip gaps relative to the blade height. These larger relative tip clearances would negate some of the cycle improvements, and ways to mitigate this eect must be found. A novel split tip blade geometry has been created which helps improve the eciency at large clearances while also improving operating range. Two identical blades are leaned in opposite directions starting at 85% span. They are cut at mid chord and the 2 halves then merged together so a split tip is created. The result is similar to the alula feathers on a soaring bird. The concept is that the split tip will energize the tip ow and increase range. For higher relative tip clearance, this will also improve eciency. The 6th rotor of a highly loaded 10 stage machine was chosen as the baseline for this study. Three dimensional CFD simulations were performed using CD Adapco\’s Star-CCM+ at 5 clearances for the baseline and split tip geometry. The choking ow and stall margin of the split tip blade was higher than that of the baseline blade for all tip clearances. The pressure ratio of the novel blade was higher than that of the baseline blade near choke, but closer to stall it decreased. The sensitivity of peak eciency to clearance was improved. At tight clearances of 0.62% of blade height, the maximum eciency of the new design was less than the baseline blade, but as the tip clearance was increased above 2.5%, the maximum eciency increased. Structural analysis was also performed to ascertain the feasibility of the design.

Secondary flows in turbomachinery and similar engineering applications are often dominated by a single streamwise vortex structure. Investigations into the control of these flows using periodic forcing have shown a discrete range of forcing frequency where the vortex is particularly receptive. Forcing in this frequency range results in increased movement of the vortex and decreased total pressure losses. Based on the hypothesis that this occurs due to a linear instability associated with the Crow instability, a fundamental study of instabilities in streamwise vortex-wall interactions is performed. In the first part of this study a three-dimensional vortex-wall interaction is computed and analyzed for the presence of convective instabilities. It is shown that the Crow instability and a range of elliptic instabilities exist in a similar form as to what has been studied in counter-rotating vortex pairs. The Crow instability is particularly affected by the presence of a solid no-slip wall. Differences in the amplification rate, oscillation angle, Reynolds number sensitivity, and transient growth are each discussed. The spatial development of the vortex-wall interaction is shown to have a further stabilizing effect on the Crow instability due to a “lift-off” behavior. Despite these discoveries, it is still shown that amplitude growth on the order of 20% is possible and transient growth mechanisms might result in an order-of-magnitude of further growth if properly initiated. With these results in mind, an experiment is developed to isolate the streamwise vortex-wall interaction. Through the use of a vortex generating wing section and a suspended splitter plate, a stable interaction is created that agrees favorably in structure to the three-dimensional computations. A small synthetic jet actuator is mounted on the splitter plate below the vortex. Phase-locked stereo-PIV velocity data and surface pressure taps both show spatial amplification of the disturbance in a frequency range which agrees well with the prediction for the Crow instability. An analysis of the vortex response shows a primarily horizontal oscillation of the vortex column which strongly interacts with the secondary vortex structure that develops in the boundary layer.

This research developed two innovative test methods that were used to experimentally evaluate the performance of a novel blade tip timing (BTT) system from Prime Photonics, LC. The research focused on creating known blade tip offsets and tip vibrations so that the results from a BTT system can be validated. The topic of validation is important to the BTT field as the results between many commercial systems still are not consistent. While the system that was tested is still in development and final validation is not complete, the blade tip offset and vibration frequency validation results show that this BTT system will be a valuable addition to turbomachinery research and development programs once completed. For the first test method custom rotors were created with specified blade tip offsets. For the blade tip offset alternate measurement, the rotors were optically scanned and analyzed in CAD software with a tip location uncertainty of 0.1 mm. The BTT system agreed with the scanned results to within 0.13 mm. Tests were also conducted to ensure that the BTT system identified and indexed the blades properly. The second developed test method used an instrumented piezoelectric blade to create known dynamic deflections. The active vibration rotor was able to create measureable deflection over a range of frequencies centered on the first bending mode of the blade. The results for the 110 Hz, 150 Hz, 180 Hz first bending resonance, 200 Hz, and 1036 Hz second bending resonance cases are presented. A strain gage and piezoelectric sensor were attached to the active blade during the dynamic deflection tests to provide an alternate method for determining blade vibration frequency. The BTT system correctly identified the active blade excitation frequencies as well as a 120 Hz frequency from the drive motor. This thesis also explored applying BTT methods and testing to more realistic blade geometry and vibration. Blade vibrations are usually classified by their frequency relative to the rotation speed. Synchronous vibrations are integer multiples of the rotational speed and are often excited by struts or vanes fixed to the engine case. For this reason, special probe placement algorithms were explored that use sine curve fitting to optimize the probe placement. Knowing how the blade will vibrate at operation before testing is critical as well. In preparation for future research, ANSYS Mechanical was used to predict the first three modes of a PT6A-28 first stage rotor blade at 1,966, 5,539, and 7,144 Hz. These frequencies were validated to within 4% using scanning laser vibrometry. The simulation was repeated at speed to produce a Campbell Diagram to highlight synchronous excitation crossings.