Detailed heat transfer measurements using transient liquid crystal thermography were performed on a novel cooling design covering the mid-chord and trailing edge region of a typical gas turbine blade under stationary and rotating conditions. The test section comprised of two channels with aspect ratio (AR) of 2:1 (mid-chord) and 4:1 (trailing edge), where the coolant was fed into the AR = 2:1 channel from the root. Rib turbulators with a pitch-to-rib height ratio (p/e) of 10 and rib height-to-channel hydraulic diameter ratio (e/D h ) of 0.075 were placed in the AR = 2:1 channel at an angle of 60° relative to the direction of flow. The coolant after entering this section was routed to the AR = 4:1 section through a set of crossover jets. The purpose of the crossover jets was to induce sideways impingement onto the pin fins that were placed in the 4:1 section to enhance heat transfer. The 4:1 section had a realistic trapezoidal shape that mimics the trailing edge of an actual gas turbine blade. The pin fins were arranged in a staggered array with a center-to-center spacing of 2.5 times the pin diameter in both spanwise and streamwise directions. The trailing edge section consisted of both radial and cutback exit holes for flow exit. Experiments were performed for a Reynolds number (Re D h (AR = 2:1) ) of 20,000 at Rotation numbers (Ro D h (AR = 2:1) ) of 0, 0.1 and 0.14. The channel averaged heat transfer coefficient on trailing side was ∼28% (AR = 2:1) and ∼7.6% (AR = 4:1) higher than the leading side for Rotation number (Ro) of 0.1. It is shown that the combination of crossover jets and pin-fins can be an effective method for cooling wedge shaped trailing edge channels over axial cooling flow designs.

Thrust collars (TCs) are bearing elements used in geared machinery that transmit axial loads from one shaft to another. TCs are primarily used in integrally geared compressors (IGCs), but are also found in gearboxes and marine propulsion applications. TCs are hydrodynamic elements featuring a converging-diverging wedge to generate a pressure field that reacts axial loads. Accurate modeling requires knowledge of the film characteristics such as cavitation, turbulence, and air ingestion, all of which reduce load capacity. Current models in the literature do not include mass-conserving cavitation algorithms or turbulent flow. The following paper introduces a new test rig that optically characterizes the thin film region of a thrust collar. The test rig geometries, speeds, and loads match those typically seen in IGC applications. The test rig utilizes a transparent acrylic window in conjunction with a high-speed camera to obtain high-speed images of the oil film. Images are filtered and averaged to obtain areas of interest in the oil film. Cavitation and turbulence areas are captured for pinion speeds of 2.5, 5, and 7.5 krpm, and axial loads of 0.5, 1, and 1.5 kN. Cavitation occurs in the diverging (upper) region of the TC and appears at pinion speeds over 5,000 rpm, but does not change in shape after that speed. The cavitation is independent of applied load. Turbulence at the inlet region (bottom) occurs at all speeds, but increases to almost 35% of the total area at the highest speed. This paper also presents a finite element (FE) model that includes predictions for the static characteristics of the TC, specifically the cavitation area. The cavitation modeling uses an iterative Elord’s method, which conserves mass. The model predicts a similar cavitation area for all speeds and loads. A computation fluid dynamics (CFD) study predicts a similar cavitation area, and pressure field to the FE model. The CFD model predicts turbulence in the lower region that increases for increasing spin speed, which matches the experimental results. The CFD model tends to underpredict the turbulence area when compared to the experiments. As IGCs move into novel application areas to satisfy new needs, the increase in efficiency and capacity comes at a cost of more load and higher speed requirements on the TCs. This work will help original equipment manufacturers model TCs more accurately to ensure safe and efficient operation.

Underplatform Dampers are commonly used in aircraft engines to limit the risk of High-Cycle Fatigue of turbine blades. The latter is located in a groove between two consecutive blades. The dry friction contact interface between the damper and the blades dissipates energy and so reduces the vibration amplitudes. Two common geometries of dampers are used nowadays, namely wedge and cylindrical dampers, but their efficiency is limited when the blades have an in-phase motion (or a motion close to it), since the damper tends to have a pure rolling motion. The objective of the present study is to analyse a new damper geometry, based on a conical shape, which prevents from this pure rolling motion of the damper and ensures a high kinematic slip. The objective of this study is to demonstrate the damping efficiency of this geometry. Hence, in a first part, the kinematic slip is approximated with analytical considerations. Then, a nonlinear dynamic analysis is performed, and the damping efficiency of this new geometry is compared to the wedge and the cylindrical geometries. The results demonstrate that the conical damper has a high damping capacity and is more efficient and more robust than the two others.

The present paper focuses on the resilience of 3-hole pressure probes to hot sand fouling in turbomachinery environments. These probes are utilized inside jet engine hot sections for diagnostics and flow characterization. Ingestion of sand and other particulates pose a significant risk to hot section components and measurement devices in gas turbine engines. In this study, wedge, cylindrical, and trapezoidal probes were exposed to hot section turbine aerothermal conditions of 1050°C and 65–70 m/s flow velocity and fouled with 0–5 μm Arizona Road Dust (ARD). Sand accumulated more rapidly on the surface of the trapezoidal and cylindrical probe geometries than on the surface of the wedge probe geometry. Probe calibrations following sand fouling were performed in an ambient temperature, open air, calibration jet at Mach 0.3 and 0.5. Calibration curves using non-dimensional coefficients were used to assess probe error in yaw angle due to sand fouling. Probe error was based on each probe’s ability to accurately measure flow direction over a flow angle range of [−10°, 10°]. On average, the probes displayed greater error at Mach 0.5 than Mach 0.3. The wedge probe performed the best after sand fouling and displayed a maximum error of less than ±2° in yaw angle. In contrast, the cylindrical probe performed the worst after sand fouling and displayed maximum errors of more than ±8° in yaw angle. Transient response did not change notably with sand fouling.

The Honeywell Uncertified Research Engine (HURE), a research version of a turbofan engine that never entered production, was tested in the NASA Propulsion System Laboratory (PSL), an altitude test facility at the NASA Glenn Research Center. The PSL is a facility that is equipped with water spray bars capable of producing an ice cloud consisting of ice particles, having a controlled particle diameter and concentration in the air flow. To develop the test matrix of the HURE, numerical analysis of flow and ice particle thermodynamics was performed on the compression system of the turbofan engine to predict operating conditions that could potentially result in a risk of ice accretion due to ice crystal ingestion. The goal of the test matrix was to provide operating conditions such that ice would accrete in either the fan-stator through the inlet guide vane region of the compression system or within the first stator of the high pressure compressor. The predictive analyses were performed with the mean line compressor flow modeling code (COMDES-MELT) which includes an ice particle model. The HURE engine was tested in PSL with the ice cloud over the range of operating conditions of altitude, ambient temperature, simulated flight Mach number, and fan speed with guidance from the analytical predictions. The engine was fitted with video cameras at strategic locations within the engine compression system flow path where ice was predicted to accrete, in order to visually confirm ice accretion when it occurred. In addition, traditional compressor instrumentation such as total pressure and temperature probes, static pressure taps, and metal temperature thermocouples were installed in targeted areas where the risk of ice accretion was expected. The current research focuses on the analysis of the data that was obtained after testing the HURE engine in PSL with ice crystal ingestion. The computational method (COMDES-MELT) was enhanced by computing key parameters through the fan-stator at multiple span wise locations, in order to increase the fidelity with the current mean-line method. The Icing Wedge static wet bulb temperature thresholds were applicable for determining the risk of ice accretion in the fan-stator, which is thought to be an adiabatic region. At some operating conditions near the splitter-lip region, other sources of heat (non-adiabatic walls) were suspected to be the cause of accretion, and the Icing Wedge was not applicable to predict accretion at that location. A simple order-of-magnitude heat transfer model was implemented into the COMDES-MELT code to estimate the wall temperature minimum and maximum thresholds that support ice accretion, as observed by video confirmation. The results from this model spanned the range of wall temperatures measured on a previous engine that experienced ice accretion at certain operating conditions. The goal of this study is to show that the computational process developed on earlier engine icing tests can be used to provide an icing risk assessment in adiabatic regions for other engines.

Modern high pressure turbine (HPT) blade design stands out due to its high complexity comprising three-dimensional blade features, multi-passage cooling system (MPCS) and film cooling to allow for progressive thermodynamic process parameters. During the last decade, probabilistic design approaches have become increasingly important in turbomachinery to incorporate uncertainties such as geometric variations caused by manufacturing scatter and deterioration. Within this scope, the first part of this two-part paper introduces parametric models for cooled turbine blades that enable probabilistic FE analysis taking geometric variability into account to aim at sensitivity and robustness evaluation. The statistical database is represented by a population of more than 400 blades whose external geometry is captured by optical measurement techniques and 34 blades that are digitized by computed tomography (CT) to record the internal geometry and the associated variability, respectively. Based on this data, parametric models for airfoil, profiled endwall (PEW), wedge surface (WSF) and MPCS are presented. The parametric airfoil model which is based on traditional profile theory is briefly described. In this regard, a methodology is presented that enables to adapt this airfoil model to a given population of blades by means of Monte-Carlo based optimization. The endwall variability of hub and shroud are parametrized by radial offsets that are applied to the respective median endwall geometry. WSFs are analytically represented by planes. Variations of the MPCS are quantified based on the radial distribution of cooling passage centroids. Thus, an individual MPCS can be replicated by applying adapted displacement functions to the core passage centroids. For each feature that is considered within the present study, the accuracy of the parametric model is discussed with respect to the variability that is present in the investigated blade population and the measurement uncertainty. Within the scope of the second part of this paper (cf. Högner et al. [1]), the parametric models are used for a comprehensive statistical analysis to reveal the parameter correlation structure and probability density functions (PDFs). This is required for the subsequent probabilistic finite element analysis involving real geometry effects.

Modern high pressure turbine (HPT) blade design stands out due to high complexity comprising three-dimensional blade features, multi-passage cooling system (MPCS) and film cooling to allow for progressive thermodynamic process parameters. During the last decade, probabilistic design approaches have become increasingly important in turbomachinery to incorporate uncertainties such as geometric variations caused by manufacturing scatter. In part B of this two-part paper, real geometry effects are considered within a probabilistic finite element (FE) analysis that aims at sensitivity evaluation. The knowledge about the geometric variability is derived based on a blade population of more than 400 individuals by means of parametric models that are introduced in part A (cf. Högner et al. [1]). The HPT blade population is statistically assessed which allows for reliable sensitivity analysis and robustness evaluation taking the variability of the airfoil, profiled endwalls (PEW) at hub and shroud, wedge surfaces (WSF) and the MPCS into account. The probabilistic method — Monte-Carlo simulation (MCS) using an extended Latin Hypercube Sampling (eLHS) technique — is presented subsequently. Afterwards, the FE model that involves thermal, linear-elastic stress and creep analysis is described briefly. Based on this, the fully automated process chain involving CAD model creation, FE mesh morphing, FE analysis and post-processing is executed. Here, the mesh morphing process is presented involving a discussion of the mesh quality. The process robustness is assessed and quantified referring to the impact on input parameter correlation. Finally, the result quantities of the probabilistic FE simulation are evaluated in terms of sensitivities. For this purpose, regions of interest are determined, wherein the statistical analysis is conducted to achieve the sensitivity ranking. A significant influence of the considered geometric uncertainties onto mechanical output quantities is observed which motivates to incorporate these in modern design strategies or robust optimization.

Radial-outflow turbo-expanders have emerged in the recent years to suit some special applications where complex multi-phase and multi-component flows (like liquid-rich gases and solid particle-laden gases) need to be expanded. This paper presents a systematic study on the design and flow behavior of a single stage radial-outflow turbo-expander, which is to be used in organic Rankine power system to covert the low-temperature heat into shaft power. A mean-line code coupled with the optimization algorithm is developed and used to carry out the one-dimensional preliminary design, where 7 non-dimensional parameters are used as design variables (nozzle velocity coefficient, rotor velocity coefficient, reaction, rotor inlet and outlet flow angles, velocity ratio, and rotor diameter ratio). In comparison with the original design, significant design performance gains are achieved with the matched combination of design parameters. Geometric shape design is further performed for the expander. In consideration of the flow features in nozzle and rotor blade passages being nearly two-dimensional, blade shape design of both rows is conducted on the basis of the airfoils used for conventional axial flow turbines, where a conformal mapping method is used to convert the axial profile into the polar coordinate frame and it is then represented by 11 parameters of mean cylindrical diameter, radial and tangential chord, leading and trailing edge radius, blade inlet and outlet angles, blade inlet wedge angle, number of blades, unguided turning, and throat size. Flow and overall performance are simulated and predicted for the designed expander, where the output shaft power and overall isentropic efficiency is respectively predicted as 87.86 kW and 81.60% at design condition.

Three-dimensional, compressible, unsteady Navier-Stokes equations are solved to investigate the unsteady flow behavior in a transonic centrifugal compressor. The computational model is a high compression ratio centrifugal compressor (4:1) consisted of an inlet duct, an impeller (15 main blades and 15 splitters) and a diffuser vane with 24 two-dimensional wedge vanes. The aim of this study is to conduct a comprehensive assessment of the ability of a hybrid scale-adaptive simulation (SAS) turbulent model to characterize the transient flow structures within the compressor passages. The main idea of SAS approach, an improved URANS (unsteady Reynold-averaged Navier-Stokes) model, is based on the introduction of von Karman length scale into the turbulent scale equation which results in LES-like behavior in unsteady regions of the flow field. A numerical sensitivity test is performed to validate the computational results in terms of pressure ratio and compressor efficiency. Instantaneous and mean flow field analyses are presented in the impeller and the vaned diffuser. Applying transient simulations, it is shown that the interaction between the pressure waves and the surface pressure of the diffuser blades leads to a pulsating behavior within the diffuser. Moreover, spectral analysis is evaluated to analyze the BPF tonal noise as the main noise source of centrifugal compressors. In addition, the current SAS results are compared with those of the URANS-SST (shear stress transport) approach to show the ability of SAS approach in the prediction of the turbulent structures where the SAS model leads to a much better resolution of the unsteady fluctuations. This study shows that the current SAS approach, as an alternative to the existing hybrid RANS/LES methods, is promising in terms of prediction of transient phenomena like LES, but with a substantially reduced turn-around time.

Ceramic matrix composite (CMC) materials are targeted for high temperature application in aircraft and power turbines, because of their low density and high-temperature thermo-mechanical properties, compared to conventional nickel super alloys. New test methods are needed for the assessment of the effects of delamination cracks on the structural integrity and life of CMC components. The ASTM C28 Fracture Toughness (Crack Growth Resistance – CGR) Working Group has drafted a standard test method for the “Mode I Interlaminar Fracture Tougness (G Ic – Crack Growth Resistance) of Fiber-Reinforced Ceramic Matrix Composites (CMC) by Wedge Loading of a Double Cantilever Beam at Ambient Temperatures” The wedge loading method was developed to avoid the problems of high temperature bonding of loading blocks and hinges. The ASTM test standard details the scope, use, and application of the test method, interferences, test equipment, specimen geometry and preparation, test procedures, data interpretation and calculation, and reporting requirements for the new CMC CGR test method.

A Durability and Damage Tolerance (D&DT) analysis of an S200 Nicalon/SiNC and Oxide/Oxide Ceramic Matrix Composite (CMC) was conducted to determine the crack growth resistance (GIc) of Wedge Loaded DCB (WDCB) at Room and Elevated temperatures (RT/ET) and compared with experimental tests observations. Wedge Loading gives proper crack path without mixed mode effects and can be used at high temperature in a furnace. Load displacement, GIc, electrical resistivity and acoustic emission was measured by tests and compared to FE based Multi Scale Progressive Failure Analysis (PFA) of the WDCB specimen. The critical damage events studied included damage initiation, damage propagation, fracture initiation, and fracture propagation as the components were being loaded. Effect of defects on Modulus (E11, E22, and E33) was conducted by Electrical Resistance (ER) Measurement at Room temperature (RT). Multi-Scale modeling simulation considered de-homogenized nano-assisted micromechanics analytical formulation, a Mori Tanaka based stiffness correction including void shape, size, distribution and orientation effects. Emitted/received signal amplitude by ER Vs. time was used to evaluate reduction of stiffness in all directions resulting in anisotropic stiffness of As-Built specimens. WDCB specimen was tested to failure at RT/ET to produce reliable GIc values with minimum specimen size. Many parameters that contribute to specimen failure included interface coating thickness, mixed mode failure evolution, interlaminar defects, delamination damage, crack bridging, and fiber fracture which were all studied in detail in this work. All simulations correlated well with test.

Oil-free micro turbomachinery (OFT) implements compliant foil bearings because of their minute drag and ability to operate in extreme (high or low) temperature. Prominent to date, bump-type foil bearings integrate an underspring thin metal structure that provides resilience and material damping; and while the rotor is airborne, acts in series with the stiffness and damping of the gas film. The design and manufacturing of foil bearings remains costly as it demands of extensive engineering and actual experience. Alternative foil bearing configurations, less costly and easier to manufacture, are highly desirable to enable widespread usage of OFT. This manuscript details the design and manufacturing of a novel Rayleigh-step metal mesh foil thrust bearing (MMFTB) as well as its testing on a dedicated rig. Metal mesh structures offer significant material structural damping and can be easily procured at a fraction of the cost of a typical bump foil strip layer. The MMFTB consists of a solid carrier, a number of stacked annular Copper mesh sheets (wire diameter = 0.25, 0.3 and 0.41 mm), and a steel top foil (0.127 mm thick) that makes six pads (ID = 50.8 mm, OD = 2 ID), each 45° in extent. A 3 μm polymer coats each pad and a photo-chemical process etches a step 20 μm in height. Static and dynamic load measurements (without rotor speed) demonstrate the MMFTB has structural stiffness and material damping similar to that of a publicized bump-type foil thrust bearing. A maiden test of the MMFTB with rotor speed of Ω = 15 krpm (∼80 m/s at bearing outer diameter) proved briefly the bearing operation when applying a tiny thrust load. Further tests with ambient air, a rotor speed of 40 krpm (∼212 m/s at bearing OD), and a very light load/area < 7 kPa failed several of the prototype bearings, all exhibiting significant wear on one or more pads. The source of the failure is the inherent unevenness of the metal mesh stacked substructures, thus causing the pads to bulge towards the rotor collar surface before a load applies. A deficient anchoring method exacerbates the unevenness. Incidentally, a high rotor speed induced large windage that lifted the top foils pushing them against the spinning collar. As the bearing moved towards the rotating collar to begin applying thrust, the local high spots rubbed against the collar, before a hydrodynamic wedge could form to separate the surfaces. Without a robust sacrificial coating, metal-to-metal contact quickly disfigured the contacting top foil pads, erasing the etched step, and leading to failure. In concept, and on paper, the mesh sheets and the top foil lay flat against the bearing carrier, giving a false sense of uniformity in the design process. In actuality, a designer must consider the manufactured states of the individual components and how they assemble. A redesign of the bearing intends to overcome the existing flaws (highlighted herein) by incorporating a thicker top foil that is well anchored (to better withstand the effects of windage), a robust sacrificial coating, and a hydrodynamic wedge accomplished via a circumferential taper on each pad.

This paper presents a numerical investigation on a concept for enhancing the film cooling performance by modifying the shape of upstream ramps. The novel shape ramp, which is placed in front of the film cooling holes, is presented to alter the approaching boundary-layer flow and its interaction with coolant to increase the lateral spreading of the coolant. Five different shape ramps are investigated, including rectangular, wedge-shaped, convex, concave and wave-shaped ramps. The film cooling performance of different shape ramps are evaluated at the density ratio about 1, with blowing ratios ranging from 0.3 to 1.2. The numerical results for the upstream ramp show an agreement with experiment data when solving three dimensional average Navier-Stokes analysis with the k-ε model. Detailed adiabatic cooling effectiveness and total pressure loss coefficient are simulated. Results obtained indicate that film cooling characteristics in the region downstream of the film cooling holes are sensitive to the ramp shapes. The wave-shaped ramp shows the lowest total pressure loss coefficient among these five ramps. For M = 0.3 and 0.6, the highest centerline adiabatic cooling effectiveness occurs in the convex ramp. And this shape ramp also shows the highest spanwise averaged adiabatic cooling effectiveness at the blowing ratio of 0.3. Compared with the other shape ramps, the concave ramp can greatly increase both the centerline and the spanwise averaged adiabatic cooling effectiveness for M = 1.0 and 1.2.

Heat transfer in a rotating two-pass trapezium-shaped channel, with staggered 90-deg ribs on both leading and trailing surfaces is experimentally investigated. The hydraulic diameter of the first and second pass is 24.5 mm and 16.9 mm, respectively. The inlet Reynolds number and rotational speed range from 10000 to 50000 and zero to 1000 rpm, respectively, which results in the inlet rotation number varying from zero to 1.0. The heated copper plate technique is employed to measure the regional averaged heater transfer coefficients. Pressure drops are measured by newly designed rotating pressure measurements module. Both ribbed cases and smooth cases are compared to present rib enhancement effect. For non-rotating result, the results show that the trailing surface presents much higher heat transfer than other cases due to the special wedge-shaped geometry. The ribbed wedge-shaped achieves enhanced regional heat transfer performances than the smooth case at all locations. Compared with the non-rotating results in the first pass, heat transfer on both trailing and leading surfaces is enhanced except for the position near the turn region, but weakened on outer surface in stream-wise direction. And at high rotation numbers, the highest maximum heat transfer on railing surface happens at a location of approximately X/D = 10. In the first pass, rotation always enhances heat transfer on the trailing surface as rotation number increases and the rotation-to-stationary Nusselt number ratio reaches to 2.0 at the rotation number of 0.5. The leading and outer surfaces both have a critical rotation number located at Ro c = 0.05.

Gas foil bearings use ambient air as working fluid, which forms the hydrodynamic wedge between the rotor surface and the compliant foil structure, therefore they are oil-free and environmental friendly. Compared to traditional bearings, gas foil bearing can operate in extreme conditions such as high temperature and high rotating speed. They also provide better damping and stability characteristics and have larger tolerance to debris and rotor misalignment. Gas foil bearings have been successfully applied to micro and small sized turbomachinery, such as micro gas turbine and cryogenic turbo expander. In the previous decades, a lot of theoretical and experimental work have been conducted to investigate the properties of gas foil bearings. However, very little work has been done to study the influence of the foil bearing pad geometry configuration using computational fluid dynamics (CFD). This study proposes a robust approach to analyze the effect of the foil geometry on the performance of a six pads thrust foil bearing. In this study, a 3D CFD model for a parallel six pads thrust foil bearing is built using ANSYS-CFX software. The full Navier-Stokes equations are solved in the fluid domain. Taking into account the rotational periodicity, the computational model is simplified and only one pad of the thrust foil bearing is analyzed. In order to predict the thermal property, the total energy with viscous dissipation is used. Practical boundary conditions are used in the baseline CFD model. The results of the baseline modes is further compared with a modified Reynolds equation solution. Based on this model, the geometry of the thrust foil bearing is parameterized and analyzed using the design of experiments (DOE) methodology. In this paper, the selected geometry parameters of the foil structure include: minimum film thickness, inlet film thickness, the ramp extent on the inner circle, the ramp extent on the outer circle, the arc extent of the pad, and the orientation of the leading edge. A factorial design technique is employed to sample the design space in DOE. The objectives in the sensitivity study are selected as load force and friction torque. An optimal foil geometry is derived based on the results of the DOE process by using a goal driven optimization technique to maximize the load force and minimize the friction torque. The results show that the geometry of foil structure is a key factor for foil bearing performance. The numerical approach proposed in this study is expected to be useful from the thrust foil bearing design perspective.

The three-dimensional, compressible, unsteady Navier-Stokes equations are solved to investigate the influence of the inclined leading edge diffuser vanes on the flow field and radiated noise from a transonic centrifugal compressor with high compression ratio. The computational domain is consisted of an inlet duct and a rotating impeller with splitter blades followed by a two-dimensional wedge vaned diffuser. The numerical method was validated by comparing the steady computational results with those of experiments in terms of pressure ratio and compressor efficiency at different operating points for the original diffuser. The transient simulations were verified by comparison of the velocity distribution with PIV data in normal flow condition before the onset of surge. In the case of steady simulations, seven types of diffuser vane with various inclination angles of leading edge were numerically modeled to investigate the effects of inclined leading edge on the diffuser pressure recovery and total pressure loss characteristics. The vaned diffuser with inclined leading edge reduces the interaction between the impeller discharge flow and diffuser leading edge which leads to improve the pressure recovery characteristics within the diffuser passage. Detailed flow analysis inside the diffuser passage showed the pressure ratio and compressor efficiency have been improved by the inclined leading edges. The maximum diffuser pressure recovery coefficient, 0.7185, and compressor efficiency, 84.80%, were observed in the case of 30 degree inclination angle from hub-to-shroud. In the case of transient simulations, five different inclined leading edge diffuser vanes were numerically conducted. The present study focuses on the unsteady pressure fluctuations and noise prediction within the impeller and diffuser passages at the compressor design point. The influences of inclination angle of diffuser vane leading edges on the pressure waves with different convective velocities, generated by the impeller-diffuser interaction and pseudo-periodic unsteady separation bubbles, were captured in the time/space domain along the diffuser blade surfaces. Since it is important to understand that the far-field acoustics are dominated by the internal pressure fluctuations inside the passages, the near-field pressure fluctuation spectra captured at the impeller-diffuser interface are evaluated to analyze the tonal BPF noise as the main noise source in the centrifugal compressors. It is shown that the inclined leading edges are very useful not only for improvement of the pressure recovery characteristics within the diffuser but also for the reduction of the interaction tonal BPF noise (around 7.6 dB SPL reduction). Furthermore, it was found that by using the inclined leading edge, the vortical structures and separations within the diffuser passages were reduced which may cause the attenuation of the broadband noise components and the overall sound pressure level.

The main focus of this study is to apply a computational tool for the flow analysis of the turbine engine that has been tested with ice crystal ingestion in the Propulsion Systems Laboratory (PSL) at NASA Glenn Research Center. The PSL has been used to test a highly instrumented Honeywell ALF502R-5 (LF11) turbofan engine at simulated altitude operating conditions. Test data analysis with an engine cycle code and a compressor flow code was conducted to determine the values of key icing parameters that can indicate the risk of ice accretion, which can lead to engine rollback (un-commanded loss of engine thrust). The full engine aerothermodynamic performance was modeled with the Honeywell Customer Deck specifically created for the ALF502R-5 engine. The mean-line compressor flow analysis code, which includes a code that models the state of the ice crystal, was used to model the air flow through the fan-core and low pressure compressor. The results of the compressor flow analyses included calculations of the ice-water flow rate to air flow rate ratio ( IWAR ), the local static wet bulb temperature, and the particle melt ratio throughout the flow field. It was found that the assumed particle size had a large effect on the particle melt ratio, and on the local wet bulb temperature. In this study the particle size was varied parametrically to produce a non-zero calculated melt ratio in the exit guide vane (EGV) region of the low pressure compressor (LPC) for the data points that experienced a growth of blockage, and resulted in an engine called rollback (CRB). At data points where the engine experienced a CRB having the lowest wet bulb temperature of 492 R at the EGV trailing edge, the smallest particle size that produced a non-zero melt ratio (between 3%–4%) was on the order of 1 μm . The particle size was varied from 1 μm – 9.5 μm to achieve the target melt ratio. For data points that did not experience a CRB which had static wet bulb temperatures in the EGV region below 492 R, a non-zero melt ratio could not be achieved even with a 1 μm ice particle size. The highest value of static wet bulb temperature for data points that experienced engine CRB was 498 R with a particle size of 9.5 μm . Based on this study of the LF11 engine test data, the range of static wet bulb temperature at the EGV exit for engine CRB was in the narrow range of 492 R – 498 R, while the minimum value of IWAR was 0.002. The rate of blockage growth due to ice accretion and boundary layer growth was estimated by scaling from a known blockage growth rate that was determined in a previous study. These results obtained from the LF11 engine analysis formed the basis of a unique icing wedge which defines a region of ice accretion risk that are being applied to other turbofan engines in order to predict the risk of ice accretion at various altitudes and operating conditions.

A Durability and Damage Tolerance (D&DT) analysis of an S200 Nicalon/SiNC CVI SiC/SiC attachment joint was conducted to determine the CMC components material structural integrity during service loading. A building block validation strategy is proposed that includes: (i) Room, and High Temperature (RT/HT) testing with AE (Acoustic Emission) and ER (Electrical Resistivity) strategies; (ii) Advanced multi-scale modeling; (iii) Interpretation test/model; and (iv) ASTM draft submittal of simplified beam equation supported by FEM/test compliance and round robin exercise. The following building block calibration, verification, validation, and accreditation strategy were performed: 1 ) Material characterization analysis to determine the damage evolution under uniaxial tensile loads and compared with test; 2 ) Crack Growth Resistance (CGR) analysis and test of wedge loaded DCB (Double Cantilever Beam) to determine the crack growth length, zig-zag pattern, fracture, shift in failure mechanisms and derivation of fracture energy vs. crack length simple formulation at RT; 3 ) Joint loading multi scale modeling and comparison with observed test load displacement curve, and determination of fracture energy; and 4 ) blade structural integrity and response under service loading using Multi-Scale Progressive Failure Analysis (MS-PFA) and determination of contributing damage and delamination types and their locations. FE based MS-PFA of the material and structure studied addressed the critical damage events (damage initiation, damage propagation, fracture initiation, and fracture propagation) as the components were being loaded. All dehomogenized multi scale methods CMC parameters were implemented in the material and structural modeling strategy, such as crack density effects and architecture (2D, 3D orthogonal, and mixed) interphase thickness, and interfacial shear strength. Many parameters that contribute to specimen failure including interface coating thickness, mixed mode failure evolution, interlaminar defects, delamination damage, crack bridging, and fiber fracture were all studied in detail in this work. Several FE-based multi-scale modeling techniques were investigated: a) MS-PFA; b) Virtual Crack Closure Technique (VCCT); and c) integrated damage and fracture evolution methodology using combined MS-PFA and VCCT.

Interlaminar fracture properties play an important role in predicting failure of structural components for CMC materials. In engine applications, components are subject to large thermal gradients which induce interlaminar stresses. One of the main challenges in evaluating interlaminar fracture toughness at room and elevated temperatures is the development of an experimental setup that provides ease for testing and allows for in-situ monitoring of the interlaminar crack growth. Therefore, a wedge-loaded DCB testing method is developed. The method utilize electrical resistance to monitor crack growth and was applied to a woven polymer infiltrated pyrolysis (PIP) SiC/SiNC composite. Post-testing inspection was carried out using optical microscopy of polished cross-sections, showing crack morphology. It was found that crack growth rate at room temperature is double the one at 815 °C for initial tests in this composite system. Estimates of Mode I energy release rate suggests flat R-curve behavior at room temperature in comparison to rising R-curve behavior at 815 °C.

Many high-speed, rotating machines across a wide range of industrial applications depend on fluid film bearings to provide both static support of the rotor and to introduce stabilizing damping forces into the system through a developed hydrodynamic film wedge. Reduced oil supply flow rate to the bearings can cause cavitation, or a lack of a fully developed film layer, at the leading edge of the bearing pads. Reducing oil flow has the well-documented effects of higher bearing operating temperatures and decreased power losses due to shear forces. While machine efficiency may be improved with reduced lubricant flow, little experimental data on its effects on system stability and performance can be found in the literature. This study looks at overall system performance of a test rig operating under reduced oil supply flow rates by observing steady-state bearing performance indicators and baseline vibrational response of the shaft. The test rig used in this study was designed to be dynamically similar to a high-speed industrial compressor. It consists of a 1.55 m long, flexible rotor supported by two tilting pad bearings with a nominal diameter of 70 mm and a span of 1.2 m. The first bending mode is located at approximately 5,000 rpm. The tiling-pad bearings consist of five pads in a vintage, flooded bearing housing with a length to diameter ratio of 0.75, preload of 0.3, and a load-between-pad configuration. Tests were conducted over a number of operating speeds, ranging from 8,000 to 12,000 rpm, and bearing loads, while systematically reducing the oil supply flow rates provided to the bearings under each condition. For nearly all operating conditions, a low amplitude, broadband subsynchronous vibration pattern was observed in the frequency domain from approximately 0–75 Hz. When the test rig was operated at running speeds above its first bending mode, a distinctive subsynchronous peak emerged from the broadband pattern at approximately half of the running speed and at the first bending mode of the shaft. This vibration signature is often considered a classic sign of rotordynamic instability attributed to oil whip and shaft whirl phenomena. For low and moderate load conditions, the amplitude of this 0.5x subsynchronous peak increased with decreasing oil supply flow rate at all operating speeds. Under the high load condition, the subsynchronous peak was largely attenuated. A discussion on the possible sources of this subsynchronous vibration including self-excited instability and pad flutter forced vibration is provided with supporting evidence from thermoelastohydrodynamic (TEHD) bearing modeling results. Implications of reduced oil supply flow rate on system stability and operational limits are also discussed.