bluuebunny/crossref_metadata_2025_split huggingface dataset:
Modelling Droplet Heat and Mass Transfer in Aero-Engine Bearing Chambers
Authors: Arcila Dafne Gaviria , Morvan Hervé , Simmons Kathy , Ambrose Stephen , Walsh Michael , Kakimpa Bruce 
Published: November 2019

Abstract
The oil inside aeroengine bearing chambers can be found in many forms, including droplets which interact with the core airflow. The ability to model such bearing chambers computationally is desirable and thus a better understanding of the evaporation process of oil droplets is of great interest. Previous studies have analyzed the flow of isothermal droplets in bearing chambers. However, further investigation is needed into the heating of droplets in the highly rotating core region. This will enable designers to evaluate the behavior of droplets in a chamber and the likelihood that they will evaporate.
The aim of this research is to analyze the oil droplet evaporation process under aeroengine bearing chamber representative conditions. An ultimate goal is the ability to predict the oil-air heat and mass transfer in the core flow region, as well as to develop an understanding of the flow inside a droplet, and how this affects evaporation. This latter is important as it has not been studied before.
This paper presents the results of a numerical study of the evaporation process of a single droplet under bearing chamber temperature and air flow conditions. The two-phase flow is simulated using ANSYS Fluent with the volume of fluid approach and the evaporation process with the “D−square law”. First, the modelling approach is validated against previous experimental and numerical analysis of fuel droplets in an air flow with heat transfer. The simulation results were in excellent agreement with a benchmarking data set. The validated approach is then applied for investigation to smaller, bearing chamber representative droplets of an oil base stock used in jet engines.
The oil evaporation rate was quantified as well as the evolution of droplet diameter, which revealed the effect of different air velocities and temperatures on the droplet. The extent to which evaporation rate increased with air velocity and temperature is quantified. It is concluded that droplets of initial diameters less than 200μm that remain in the chamber core region for more than 0.3s are likely to evaporate completely.
This study allows us to estimate droplet heat and mass transfer and the associated phase change in a bearing chamber. It also provides best practice to predict the performance of small droplets under the effects of high temperature and velocity convective air flows. In future work this methodology will be applied in simulations in a representative bearing chamber to predict how the cooling process is affected by oil evaporation.
Emulation of As-Manufactured Transonic Rotor Airfoil Modal Behavior and the Significance of Frequency Veering
Authors: Brown Jeffrey M. , Carper Emily B. , Beck Joseph A. , Kaszynski Alexander A. 
Published: November 2019

Abstract
This effort investigates the ability to efficiently emulate the modal characteristics of a transonic compressor rotor as a function of measured geometric variations. The transonic rotor is measured with a structured blue light scanner that provides a dense geometric representation of the as-manufactured part. Principal component analysis (PCA) is applied to create a reduced order, orthogonal basis of parameters used as independent emulator parameters. Ordinary least squares (OLS) and Gaussian stochastic process (GSP) regression models are employed as emulators for airfoil frequency and mode shape across the first twenty resonant modes. While many modes are emulated with high accuracy, some challenge conventional emulation. It is shown that geometric variations cause significant variation in modal behavior for closely spaced frequencies. This work identifies that as-manufactured geometry deviations can lead to frequency veering and associated chaotic modal behavior. Effective emulation of frequncy and mode shape is demonstrated, but the physics-based understanding of as-manufactured geometry frequeny veering defines the way for future improvements in emulation accuracy and computational requirements.
Experimental Investigation of Sweeping Jet Film Cooling in a Transonic Turbine Cascade
Authors: Hossain Mohammad A. , Asar Munevver E. , Gregory James W. , Bons Jeffrey P. 
Published: November 2019

Abstract
The sweeping jet (SJ) film cooling hole has shown promising cooling performance compared to the standard shaped hole in low-speed conditions. The present work demonstrates the first attempt of sweeping jet film cooling at an engine relevant Mach number. An experimental investigation was conducted to study the sweeping jet film cooling on a nozzle guide vane suction surface. A well-established additive manufacturing technique commonly known as Stereolithography (SLA) was unitized to design a transonic, engine representative vane geometry in which a row of SJ holes was used on the vane suction surface. Experiments were performed in a linear transonic cascade at an exit Mach number of 0.8 and blowing ratios of BR = 0.25–2.23. The measurement of heat transfer was conducted with the transient IR method and the convective heat transfer coefficient (HTC) and adiabatic film cooling effectiveness were estimated using a dual linear regression technique (DLRT). Aerodynamic loss measurements were also performed with a total pressure Kiel probe at 0.25Cax downstream of the exit plane of the vane cascade. Experiments were also conducted for a baseline shaped hole (777-hole) for a direct comparison. Results showed that the SJ hole has a wider coolant spreading in the lateral direction near the hole exit due to its sweeping action that improves the overall cooling performance particularly at high blowing ratios (BR>1). Aerodynamic loss measurement suggested that the SJ hole has a comparable total pressure loss to the 777-shaped hole.
Characterisation of Waves and Ligaments in Films Close to an Aeroengine Ball Bearing
Authors: Hee Jee Loong , Simmons Kathy , Hann David , Walsh Michael 
Published: November 2019

Abstract
Surface waves are observed in many situations including natural and engineering applications. Experiments conducted at the Gas Turbine and Transmissions Research Centre (G2TRC) used high speed imaging to observe multiscale wave structures close to an aeroengine ball bearing in a test rig. The dynamic behavior and scale of the waves indicate that these are shear-driven although highly influenced by gravity at low shaft speed.
To understand the interactions between gas and liquid phases including momentum and mass transfers, characterization of the observed waves and ligaments was undertaken. Waves were studied at surfaces close to the ball bearing and ligaments were assessed near the cage. Characterization was in terms of frequency and wavelength as functions of speed, flow-rate, bearing axial load and gravity. The assessments confirmed the existence of gravity-capillary waves and capillary waves.
Gravity-capillary waves were measured to have a longer mean wavelength on the co-current side of the bearing (gravity and shear acting together) compared to the counter-current side (gravity and shear opposing). Using a published definition of critical wavelength (λcrit), measured wavelengths at 3,000 rpm were 2.56λcrit on the co-current side compared to 1.86λcrit at the countercurrent location. As shaft speed increases, wavelength reduces with transition to capillary waves occurring at around 0.83λcrit. At shaft speeds beyond 5000 rpm, capillary waves were fully visible and the wavelength was obtained as 0.435λcrit. Flow-rate and load did not significantly influence wavelength.
Wave frequency was found to be proportional to shaft speed. The gravity-capillary waves had frequencies within the range 13–25 Hz while capillary waves exhibited a frequency well beyond 100 Hz. The frequencies are highly fluctuating with no effect of load and flow rate observed.
Ligaments were characterized using Weber number and Stability number. The number of ligaments increased with shaft speed. A correlation for ligament number based on operating conditions is proposed.
A Low Reynolds Number Experimental Evaluation of Tubercles on a Low-Pressure Turbine Cascade
Authors: Pym Stephen A. , Asghar Asad , Allan William D. E. , Clark John P. 
Published: November 2019

Abstract
Aircraft are operating at increasingly high-altitudes, where decreased air density and engine power settings have led to increasingly low Reynolds numbers in the low-pressure turbine portion of modern-day aeroengines. These operating conditions, in parallel with highly-loaded blade profiles, result in non-reattaching laminar boundary layer separation along the blade suction surface, increasing loss and decreasing engine performance. This work presents an experimental investigation into the potential for integrated leading-edge tubercles to improve blade performance in this operating regime. A turn-table cascade test-section was constructed and commissioned to test a purpose-designed, forward-loaded, low-pressure turbine blade profile at various incidences and Reynolds numbers. Baseline and tubercled blades were tested at axial chord Reynolds numbers at and between 15 000 and 60 000, and angles of incidence ranging from −5° to +10°. Experimental data collection included blade surface pressure measurements, total pressure loss in the blade wakes, hot-wire anemometry, surface hot-film measurements, and surface flow visualization using tufts. Test results showed that the implementation of tubercles did not lead to a performance enhancement. However, useful conclusions were drawn regarding the ability of tubercles to generate stream-wise vortices at ultra-low Reynolds numbers. Additional observations helped to characterize the suction surface boundary layer over the highly-loaded, low-pressure turbine blade profile when at off-design conditions. Recommendations were made for future work.
An Analytical Shock-Loss Model Fully Integrated to a Streamline Curvature Method for Axial-Flow Compressors
Authors: Aguirre Hasani Azamar , Pachidis Vassilios , Templalexis Ioannis 
Published: November 2019

Abstract
The constant and increasing demand to obtain more accurate turbomachinery performance prediction in the design and analysis process has led to the development of higher fidelity flow field models. Despite extensive flow field information can be collected from 3-D RANS numerical simulations, the computational cost is expensive in terms of time and resources, especially if they are used as solvers within a design-optimisation framework. In contrast, 2-D throughflow methods, such as streamline curvature (SLC), provide an acceptable flow solution in minutes.
The use of modern and advanced-design transonic axial-flow compressors and fans has been expanding due to their high shock-induced single-stage pressure ratios while being light, compact and robust. Transonic-flow analysis in blading is complex due to the shock structures involved and associated phenomena. Previous 2-D SLC tools have failed to replicate the real compressible-flow physics, assuming and oversimplifying the shock-system shape and location. The situation aggravates, when the assumed overall shock configuration applies only for design point at unstarted operations, requiring of empirical correlations to estimate the shock-loss coefficient for off-design operations. The overall compressor performance prediction is thence highly-dependent on the shock modelling quality.
For this reason, a physics-based shock -structure and -loss model was developed and implemented into an existing in-house 2-D SLC compressor performance simulator to enhance the aerodynamic prediction in transonic axial-flow compressors. The novel shock-loss model is fully coupled to the 2-D SLC software, for which a blade-element-layout method was adapted to obtain the profile geometry definition. The analytical shock-loss model possesses the capability to operate at started and unstarted passages utilizing an iterative-solution method to position the choke-induced passage-shock. A significant contribution of the new shock-loss model is the solution of the relative total-pressure loss for the entire blade span, comprising the inlet relative subsonic supercritical and supersonic regions.
In this manner, shock losses were determined throughout the blade span and for various off-design operating conditions, including those at choking. 2-D SLC simulations were conducted for the NASA Rotor 67 Fan to validate the models accordingly against test-rig data and verify against previous model estimations and 3-D CFD results. The analytical shock - structure and -loss model improved the shock-loss prediction between 40–50% with respect of the state-of-the-art models and showed satisfactory agreement against measured data within 0.6% at the blade tip and 0.3% at mid-span sections.
Evaluation of Spray Performance of Pyrolysis Oil
Authors: Yun Sangsig , Choi Minji , Kumar Ashwani 
Published: November 2019

Abstract
Pyrolysis oil has become an important subject of research as it is considered to be a potential environmentally friendly and cheap alternative to conventional fossil fuels. Unfortunately, due to the significant differences of the chemical and physical properties of pyrolysis oil than that of fossil fuels, the deployment of pyrolysis oil in existing power systems such as gas turbines and internal combustion engines has been highly restricted. Thus, major research on pyrolysis oil has been conducted to overcome these challenges related to the unfavorable physical and chemical properties of pyrolysis oil. This paper reports experimental work on the effects of physical properties of the pyrolysis oil on spray performance of nozzles. Effort to evaluate the spray performance by using different types of atomizers has been made as well. Laser based diagnostics was applied to obtain qualitative comparisons spray characteristics of various pyrolysis oils. Experimental data such as the distribution of fuel droplet sizes and overall spray shapes was analyzed, which could provide valuable guidelines to design fuel nozzles. Lastly, the paper will also present NRC’s plans to accelerate the deployment of such pyrolysis oils in industrial gas turbines.
High-Speed Spectrally Resolved Imaging Studies of Spherically Expanding Natural Gas Flames Under Gas Turbine Operating Conditions
Authors: Parajuli Pradeep , Paschal Tyler , Turner Mattias A. , Petersen Eric L. , Kulatilaka Waruna D. 
Published: November 2019

Abstract
Natural gas is a major fuel source for many industrial and power-generation applications. The primary constituent of natural gas is methane (CH4), while smaller quantities of higher order hydrocarbons such as ethane (C2H6) and propane (C3H8) can also be present. Detailed understanding of natural gas combustion is important to obtain the highest possible combustion efficiency with minimal environmental impact in devices such as gas turbines and industrial furnaces. For a better understanding the combustion performance of natural gas, several important parameters to study are the flame temperature, heat release zone, flame front evolution, and laminar flame speed as a function of flame equivalence ratio. Spectrally and temporally resolved, high-speed chemiluminescence imaging can provide direct measurements of some of these parameters under controlled laboratory conditions. A series of experiments were performed on premixed methane/ethane-air flames at different equivalence ratios inside a closed flame speed vessel that allows the direct observation of the spherically expanding flame front. The vessel was filled with the mixtures of CH4 and C2H6 along with respective partial pressures of O2 and N2, to obtain the desired equivalence ratios at 1 atm initial pressure. A high-speed camera coupled with an image intensifier system was used to capture the chemiluminescence emitted by the excited hydroxyl (OH*) and methylidyne (CH*) radicals, which are two of the most important species present in the natural gas flames. The calculated laminar flame speeds for an 80/20 methane/ethane blend based on high-speed chemiluminescence images agreed well with the previously conducted Z-type schlieren imaging-based measurements. A high-pressure test, conducted at 5 atm initial pressure, produced wrinkles in the flame and decreased flame propagation rate. In comparison to the spherically expanding laminar flames, subsequent turbulent flame studies showed the sporadic nature of the flame resulting from multiple flame fronts that were evolved discontinuously and independently with the time. This paper documents some of the first results of quantitative spherical flame speed experiments using high-speed chemiluminescence imaging.
Adiabatic Film Cooling Effectiveness Measurement of High Performance Combustion Liner Slot Jets
Authors: Flores Wilmer , Curbelo Andres , Calderon Luisana , Kapat Jayanta S. , Ahmed Kareem 
Published: November 2019

Abstract
The present study experimentally investigates adiabatic film cooling effectiveness for high performance flow auxetic geometries (S-Slot and Ellipsoid). Both geometries are auxetic structures that were designed around a Negative Poisson Ratio (NPR). Manufactured NPR combustion liners offer higher quality fracture resistance [1] and high energy absorption [2]. Film cooling effectiveness results for test geometries were non-dimensionalized by a baseline case of circular inline orifices. Non-intrusive pressure sensitive paint (PSP) technique was utilized to obtain local surface pressure gradients for each experimental geometry. Adiabatic film cooling effectiveness measurements were tested for multiple blowing ratios (BR) ranging between 0.05–1.15. Three plate configurations were studied with void inclination angles of 0 degrees. Analyses demonstrate that blowing ratios of 0.5 and 0.75 are optimal for S-slot adiabatic effectiveness. Film cooling results for the Ellipse configuration show increasing effectiveness up to an optimal blowing ratio of 0.5, and decreases as BR is increased beyond that. Additional experiments were performed by keeping the differential pressure across the plates constant. Results indicate that the S-slot provides optimum film cooling effectiveness while the differential pressure conserved in comparison to Ellipses and Circles.
Measurements of End Cap Temperatures in a Multistage Barrel-Style Centrifugal Compressor
Authors: Smith Natalie R. , Wilkes Jason C. , Wade Jonathan , Allison Timothy C. , Day Towler Meera , Moore J. Jeffrey , Weinberg Ian , McCune Michael 
Published: November 2019

Abstract
Gas compressor customers desire to operate gas compressors at higher discharge temperatures. In barrel-style centrifugal compressors, dry gas seals are used to limit the leakage along the shaft from the process fluid to atmosphere. The temperature limit of dynamic O-ring seals in the dry gas seal cartridge is one of the primary limiters to operating gas compressors at higher discharge temperatures. While the present discharge temperature limit is conservative compared to the seal limits, without a detailed understanding of the thermal distribution around the discharge-side end cap, operation at compressor discharge temperatures that exceed seal limits has too much risk. Thus, this paper describes the approach taken to characterize the temperatures around the discharge end dry gas seal of a commercial C335EL centrifugal compressor. A phased test matrix with increasing discharge temperatures was conducted so that the temperature distribution throughout the end cap could be assessed. The full-scale, highly-instrumented compressor was operated at discharge pressures and temperatures ranging from 4.7 to 9.4 MPa (680 to 1370 psi) and 113 to 218 degree Celsius (235 to 425 degrees Fahrenheit), respectively. The experimental test set-up and results are presented herein. Results include demonstration of successful compressor operation at discharge temperatures greater than seal limits as well as the end cap temperature sensitivity to lube oil supply and dry gas supply temperatures.
Analysis of Gas Turbine Rim Cavity Ingestion With Axial Purge Flow Injection
Authors: Robak Christopher W. , Faghri Amir , Thole Karen A. 
Published: November 2019

Abstract
Turbine rim cavities require an adequate supply of cooling purge flow to prevent hot gas ingestion from overheating metal components beneath the gas path airfoils. Purge flow is typically introduced into rim cavities through a labyrinth seal at the inner diameter of the cavity, or through conduits in the metal walls of the rim cavity. This numerical study will focus on purge flow introduced through axial holes in the stationary side of a turbine realistic rim cavity. Three dimensional Unsteady Reynolds-average Navier-Stokes (URANS) CFD modeling is utilized to model of cavity sealing effectiveness as a function of axial purge flow rate. CFD modeling is compared with experimental data from the test turbine in the Steady Thermal Aero Research Turbine (START). Results show good agreement with experimental data, especially at lower purge flow rates. Analytical depictions of the flow field setup in the rim cavity are provided, explaining trends observed in experimental data. Differences in sealing effectiveness trends between the upper and lower portions of the rim cavity are predicted by CFD modeling, adding insight to ingestion phenomena in turbine realistic rim cavities with complex geometry and flow leakage paths.
Detailed Experimental Measurement and RANS Simulation of a Low Pressure Turbine With High Lift Blading
Authors: Perez Ethan , Schmitz John T. , Jaffa Nicholas A. , Jemcov Aleksandar , Cameron Joshua D. , Morris Scott C. 
Published: November 2019

Abstract
The aerodynamic characteristics of high–lift airfoil designs is of interest for improved performance and reduced blade count in Low–Pressure Turbine (LPT) design. The present paper presents both experimental measurements as well as numerical simulation results from a single-stage LPT. The airfoils were designed for an embedded stage with a total pressure expansion ratio of 1.75 and a rotor Zweifel coefficient of 1.35. The measurement program was highly unique in that detailed measurements were obtained using a variety of different probe types, including time–resolved total pressure and hot–wires. Agreement between various measurement types was generally good, but differences beyond typically stated uncertainty bounds were noted. The computations were done using RANS and a mixing model via commercially available software. The numerical results were evaluated to determine the efficacy of this type of model for prediction and design of high–lift airfoils. The computations agreed very well with the experimental results in the midspan region, but losses were over–predicted in the lower 40% span near the hub. A basic description and understanding of the flow physics in the LPT stage are presented based on the relative agreement between the experiments and computations.