Master of Science in Mechanical Engineering (MSME), Wright State University, 2024, Mechanical Engineering

In the constant search for more efficient engines, one approach to gain performance is to reduce the weight of the low pressure turbine (LPT) module. This module can account for up to 30% of the total engine weight [1], and a reduction in LPT weight results in clear gains to engine performance and a reduction in engine cost. High lift airfoils accomplish this weight reduction by each blade extracting a larger amount of work from the flow and thus requiring fewer blades to drive the compressor when compared to conventional blades. However, high lift LPT blades, quantified by a high Zweifel loading coefficient Zw>1.15, encounter increasing loss at low Reynolds numbers. Named Reynolds lapse, this effect is problematic if the engine must operate at high altitude cruise conditions such as the case with unmanned air vehicles. The two airfoils of this study, the L2FHW and the L3FHW, were designed to be front loaded and to demonstrate favorable low Reynolds number loss characteristics. Both airfoils were tested in the Transonic Turbine Cascade (TTC) at the Air Force Research Laboratory Building 18 Test Cell 21. The TTC is capable of high Mach number and low Reynolds number flow via independent control of each. Each airfoil was tested across a broad range of Mach numbers: exit Mach 0.78 down to 0.2 and exit Reynolds numbers from 23,000 to 201,000. Across each condition an exit total pressure traverse yielded the loss coefficient of the cascade at that condition. It was found that across all design exit Mach conditions, 0.78, both airfoils experience fully attached flow and nearly flat loss behavior. This strongly aligns with the design level predictions made. At conditions beyond expected operating conditions, the L2FHW displayed resistance to un-reattaching separations at all conditions down to exit Mach 0.2 Reynolds number 23,300. The L3FHW showed un-reattaching separations at only the most extreme condition tested, exit Mach 0.2 and Reynolds number 25,300. (open full item for complete abstract)

Committee: Mitch Wolff Ph.D. (Advisor); Andrew Lethander Ph.D. (Committee Member); John Clark Ph.D. (Committee Member) Subjects: Aerospace Engineering

Doctor of Philosophy, The Ohio State University, 2023, Aero/Astro Engineering

The feasibility of an active flow control enabled variable area turbine was explored. Pressurized air was ejected from the nozzle guide vanes to reduce the effective choke area, and mass flow rate through, the turbine inlet. A set of experimental and computational studies were conducted with varying actuator types and parameters to determine their effectiveness and develop models of the flow physics. Preliminary results from a simple quasi-1D converging-diverging nozzle, with an injection flow slot upstream of the throat, showed a 2.2:1 ratio between throttled mass flow rate and injected mass flow rate at a constant nozzle pressure ratio. The penetration of the injection flow and corresponding reduction in the primary flow streamtube were successfully visualized using a shadowgraph technique. Building on this success, a representative single passage nozzle guide vane transonic flowpath was constructed to demonstrate feasibility beyond the quasi-1D converging-diverging nozzle. Both secondary slot blowing from the vane pressure surface and vane suction surface just upstream of the passage throat again successfully reduced primary flow. In addition, fluidic vortex generators were used on the adjacent suction surface to reduce total pressure loss along the midspan and further throttle the primary flow. Computational fluid dynamics simulations were used to explore the effects of a variety of parameters on the flow blockage and actuator effectiveness. Simplified models were developed to describe the relationships of various factors impacting flow blockage, turning angle, and total pressure loss. Finally, the active flow control systems were simulated at engine relevant pressures and temperatures and found to have only a minimal drop in total pressure recovery and effectiveness, which could be predicted by the simplified blockage model.

Committee: Jeffrey Bons (Advisor); Datta Gaitonde (Committee Member); Randall Mathison (Committee Member) Subjects: Aerospace Engineering

Doctor of Philosophy (Ph.D.), University of Dayton, 2023, Aerospace Engineering

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.

Committee: Markus Rumpfkeil (Advisor); Christopher Marks (Committee Member); Sidaard Gunasekaran (Committee Member); John Clark (Committee Member) Subjects: Aerospace Engineering

Master of Science in Engineering, University of Akron, 2021, Engineering

This research quantifies the high-pressure rheological performance of various jet engine turbine oils and helicopter transmission oils. The jet engine oils are classified against both the MIL-PRF-23699 HTS and AS 5780 HPC lubricant performance specifications. The helicopter transmission oils are classified against the DOD-PRF-85734 lubricant performance specification. Rheological properties include the low shear viscosity as a function of temperature and pressure, apparent shear viscosity as a function of temperature, pressure, and shear stress, and lubricant relative volume as a function of temperature and pressure. The low shear viscosity was obtained with a set of pressurized falling body viscometers capable of measuring viscosity at pressures on the order of 1GPa. Apparent shear viscosity was observed using a high pressure Couette viscometer capable of subjecting fluids to 20MPa shear stress. Lastly, density was determined using a relative volume bellows which enables density measurements up to 350MPa of pressure. In addition to the rheological properties, the lubricant's maximum traction coefficients were investigated using a full film traction test. Models were regressed to describe each lubricant's rheological properties. These include the Williams, Landel, and Ferry (WLF) Modified Yasutomi viscosity model, the Modified Carreau Yasuda and Double Modified Carreau Yasuda shear viscosity models, as well as both the Tait and Murnaghan relative volume models used in describing the lubricants' compressibility. In addition to the stated models, thermal aspects of the lubricants were deduced based on scaling rules available in literature. By quantifying the rheological properties of these lubricants, in addition to their thermal characteristics, bearing and gear suppliers can utilize the advanced models presented herein to tailor internal geometries that exploit the lubricants' response to pressure, temperature, and shear stress which In turn, further reduces frictio (open full item for complete abstract)

Committee: Gary Doll (Advisor); Yalin Dong (Committee Member); Alper Buldum (Committee Member) Subjects: Aerospace Engineering; Aerospace Materials; Engineering

Master of Science, The Ohio State University, 2021, Mechanical Engineering

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 (open full item for complete abstract)

Committee: Randall Mathison PhD (Advisor); Michael Dunn PhD (Committee Member) Subjects: Aerospace Engineering; Engineering; Experiments; Mechanical Engineering

Master of Science, The Ohio State University, 2019, Mechanical Engineering

Aerodynamic losses due to tip leakage in a High Pressure Turbine (HPT) rotor setting are analyzed using the commercial CFD software Star-CCM+. The compressible RANS equations were solved over unstructured polyhedral grids of 11 to 17 million cells, with turbulence resolved using the k-w SST model. A study of 3 blade tip/casing configurations at tip clearances between 0% and 4% of blade span is performed, with resulting mass-flow Reynolds #'s of 2.4-2.7 million and levels of reaction of 0.13-0.17. A flat-tip blade experienced a linear correlation between tip clearance and efficiency. The same blade with recessed casing was more efficient for small clearances, but large shear losses were seen to nullify this advantage at high clearances. A shrouded blade was more efficient than both at all clearances, displaying asymptotic behavior at high clearances. Various pressure and velocity monitoring techniques are employed to evaluate leakage flow and loss. Location and strength of leakage vortices are discussed, as these are closely tied with efficiency. A parametric study was performed on the shroud geometry, examining seal angle, number of seals, seal wear, and shroud axial gap. Efficiencies were compared between each study point at clearances of 2% and 4% of blade span. Total pressure loss was tracked across the blade and split between the main flow path and tip gap to distinguish between Profile and Internal Gap Shear losses. Axial gap was seen to have the greatest impact on efficiency, as a narrower gap acts as additional tip flow blockage and increases efficiency. At lower clearances, seal wear produced a negative impact as well.

Committee: Jeffrey Bons Dr. (Advisor); Randall Mathison Dr. (Committee Member); Ali Ameri Dr. (Committee Member); Robert Boyle (Committee Member) Subjects: Aerospace Engineering; Mechanical Engineering

Doctor of Philosophy, The Ohio State University, 2016, Mechanical Engineering

An experiment was performed at The Ohio State University Gas Turbine Laboratory for a film-cooled high-pressure turbine stage operating at design-corrected conditions, with variable rotor and aft purge cooling flow rates. Several distinct experimental programs are combined into one experiment and their results are presented. Pressure and temperature measurements in the internal cooling passages that feed the airfoil film cooling are used as boundary conditions in a model that calculates cooling flow rates and blowing ratio out of each individual film cooling hole. The cooling holes on the suction side choke at even the lowest levels of film cooling, ejecting more than twice the coolant as the holes on the pressure side. However, the blowing ratios are very close due to the freestream massflux on the suction side also being almost twice as great. The highest local blowing ratios actually occur close to the airfoil stagnation point as a result of the low freestream massflux conditions. The choking of suction side cooling holes also results in the majority of any additional coolant added to the blade flowing out through the leading edge and pressure side rows. A second focus of this dissertation is the heat transfer on the rotor airfoil, which features uncooled blades and blades with three different shapes of film cooling hole: cylindrical, diffusing fan shape, and a new advanced shape. Shaped cooling holes have previously shown immense promise on simpler geometries, but experimental results for a rotating turbine have not previously been published in the open literature. Significant improvement from the uncooled case is observed for all shapes of cooling holes, but the improvement from the round to more advanced shapes is seen to be relatively minor. The reduction in relative effectiveness is likely due to the engine-representative secondary flow field interfering with the cooling flow mechanics in the freestream, and may also be caused by shocks and other compr (open full item for complete abstract)

Committee: Randall Mathison (Advisor); Michael Dunn (Committee Member); Sandip Mazumder (Committee Member); Jeffrey Bons (Committee Member) Subjects: Aerospace Engineering; Engineering; Mechanical Engineering

Doctor of Philosophy (PhD), Wright State University, 2007, Engineering PhD

This dissertation is a design and validation study of the high-lift low-pressure turbine (LPT) blade designated L2F. High-lift LPTs offer the promise of reducing the blade count in modern gas turbine engines. Decreasing the blade count can reduce development and maintenance costs and the weight of the engine, but care must be taken in order to maintain turbine section performance with fewer blades. For an equivalent amount of work extracted, lower blade counts increase blade loading in the LPT section. The high-lift LPT presented herein allows 38% fewer blades with a Zweifel loading coefficient of 1.59 and maintains the same inlet and outlet blade metal angles of conventional geometries in service today while providing an improved low-Reynolds number characteristic. The computational design method utilizes the Turbine Design and Analysis System (TDAAS) developed by John Clark of the Air Force Research Laboratory. TDAAS integrates several government-funded design utilities including airfoil and grid generation capability with a Reynolds-Averaged Navier-Stokes flow solver into a single, menu-driven, Matlab-based system. Transition modeling is achieved with the recently developed model of Praisner and Clark, and this study validates the use of the model for design purposes outside of the Pratt & Whitney (P&W) design system where they were created. Turbulence modeling is achieved with the Baldwin and Lomax zero-equation model. The experimental validation consists of testing the front-loaded L2F along with a previously designed, mid-loaded blade (L1M) in a linear turbine cascade in a low-speed wind tunnel over a range of Reynolds numbers at 3.3% freestream turbulence. Hot-wire anemometry and pressure measurements elucidate these comparisons, while a shear and stress sensitive film (S3F) also helps describe the flow in areas of interest. S3F can provide all 3 components of stress on a surface in a single measurement, and these tests extend the operational envelope of the (open full item for complete abstract)

Committee: Mitch Wolff (Advisor) Subjects:

MS, University of Cincinnati, 2003, Engineering : Mechanical Engineering

Low-pressure turbines (LPT) in aircraft engines undergo tremendous losses at cruise conditions. The flow Reynolds number at cruise is lower than the take-off Reynolds number by a factor of almost two. At low Reynolds numbers, the flow is largely laminar, and tends to separate easily on the suction surface of the turbine blade when an adverse pressure gradient is encountered. Therefore, accurate prediction of flow separation is crucial for an effective design of LPT blade; and is achieved in the present work using a high-order accurate numerical solution procedure. The three-dimensional, unsteady, full Navier-Stokes equations are solved to analyze the flow. A MPI-based higher-order, parallel, chimera version of the FDL3DI flow solver, is extended for use with this turbomachinery application. A sixth-order accurate compact-difference scheme is used for the spatial discretization, along with second-order accurate temporal discretization. Tenth-order filtering is used to minimize the numerical oscillations in the flow solution and maintain numerical stability. The objective of the present study is to show the ability of higher-order accurate compact-difference scheme to predict the flow separation that occurs inside an LPT cascade at Re C = 25,000 (based on axial chord and inlet velocity). A new set of subsonic inflow/outflow boundary conditions that account for upstream influence (BC2) are derived by specifying stagnation quantities at the inlet, and a static quantity at the exit of the flow domain, and maintaining the inflow angle constant. For inflow/outflow boundary conditions that do not account for upstream influence, fixed inflow with extrapolated outflow (BC1) has been utilized. The effect of the two different sets of inflow/outflow boundary conditions on the flow solution is studied, for second-order, fourth-order and sixth-order accurate schemes. The computed Cp distribution for the LPT flow shows good agreement with the existing experimental data. The locatio (open full item for complete abstract)

Committee: Dr. Urmila Ghia (Advisor) Subjects: Engineering, Mechanical

Doctor of Philosophy, The Ohio State University, 2011, Mechanical Engineering

Rotor purge flow cavity seals are used in gas turbine engines to prevent ingestion of the mainstream gas flow into the purge cavity. Ingestion into this cavity leads to an increase in the cavity air temperature and subsequently to the rotor disk and stator metal temperatures leading to higher thermal stresses and reduced disk and stator fatigue life. An over designed cavity seal with an excess amount of purge flow has the downside of increasing engine fuel consumption through reduced turbine efficiency. The opposite approach of strengthening the hardware to withstand the higher stress and temperatures would increase the weight of the propulsion system. Understanding how the purge flow cavity and cavity seals interact with the mainstream gas is important to producing a balanced design between weight, fuel consumption, efficiency, and fatigue life of surrounding hardware. The main objective of this research was to perform an experimental and computational study of a one and one half stage high-pressure turbine installed at The Ohio State University Gas Turbine Laboratory Turbine Test Facility with emphasis on the rotor purge cavity. The rig housing the turbine stage incorporated many features found in a typical commercial high-pressure turbine such as a cooled high-pressure vane row with hub and shroud cooling, a downstream blade row followed by a downstream vane row, the ability to created elevated radial inlet temperature profiles using a combustor emulator, and a cooling supply line to the purge cavity. Multiple runs were performed to study the effects of cooling flows from both an aerodynamic and heat transfer perspective and incorporated instrumentation throughout the rig in order to capture time-accurate temperature, pressure, and heat flux measurements. The run matrix included cold rig configurations with no cooling flow, high-temperature uniform inlet profiles at the vane inlet for cases with and without cooling flows, and high-temperature radial inlet profile (open full item for complete abstract)

Committee: Michael Dunn (Advisor); Sandip Mazumder (Committee Member); Mohammad Samimy (Committee Member); Jen-Ping Chen (Committee Member) Subjects: Aerospace Engineering; Mechanical Engineering

Doctor of Philosophy, The Ohio State University, 2003, Aeronautical and Astronautical Engineering

This research uses a modern 1 and 1/2 stage high-pressure (HP) turbine operating at the proper design corrected speed, pressure ratio, and gas to metal temperature ratio to generate a detailed data set containing aerodynamic, heat-transfer and aero-performance information. The data was generated using the Ohio State University Gas Turbine Laboratory Turbine Test Facility (TTF), which is a short-duration shock tunnel facility. The research program utilizes an uncooled turbine stage for which all three airfoils are heavily instrumented at multiple spans and on the HPV and LPV endwalls and HPB platform and tips. Heat-flux and pressure data are obtained using the traditional shock-tube and blowdown facility operational modes. Detailed examination show that the aerodynamic (pressure) data obtained in the blowdown mode is the same as obtained in the shock-tube mode when the corrected conditions are matched. Various experimental conditions and configurations were performed, including LPV clocking positions, off-design corrected speed conditions, pressure ratio changes, and Reynolds number changes. The main research for this dissertation is concentrated on the LPV clocking experiments, where the LPV was clocked relative to the HPV at several different passage locations and at different Reynolds numbers. Various methods were used to evaluate the effect of clocking on both the aeroperformance (efficiency) and aerodynamics (pressure loading) on the LPV, including time-resolved measurements, time-averaged measurements and stage performance measurements. A general improvement in overall efficiency of approximately 2% is demonstrated and could be observed using a variety of independent methods. Maximum efficiency is obtained when the time-average pressures are highest on the LPV, and the time-resolved data both in the time domain and frequency domain show the least amount of variation. The gain in aeroperformance is obtained by integrating over the entire airfoil as the three-dim (open full item for complete abstract)

Committee: Michael Dunn (Advisor) Subjects: Engineering, Aerospace