PhD, University of Cincinnati, 2023, Engineering and Applied Science: Aerospace Engineering

Boundary layer ingested engines have the potential to offer significantly reduced fuel burn, but the fan stage must be designed to run efficiently with a distorted inflow. It must also be able to withstand unsteady aerodynamic loads resulting from a complex non-uniform flowfield. This research work applies different numerical methods for an improved understanding of the aerodynamic interaction between a transonic fan and inlet distortion. A single stage transonic tail cone thruster fan was designed using both in-house and commercial tools operating in an inlet distortion flowfield. This paper demonstrates that the relevant metrics required to compute the aerodynamic performance of a fan stage in distorted conditions can be reasonably modeled with a few harmonics using the non-linear harmonic method in a fraction of time compared to the full annulus time marching method. The flowfield through a transonic fan/compressor cascade shows a very complex structure due to the presence of shock waves and bladerow trailing edge wakes. In addition, the interaction of these shock waves with the blade boundary layer inherently leads to a very complex flow behavior. The second part of this investigation quantified this behavior and its influence on the stage performance and described the occurring transonic flow phenomena in detail. The last part of this research focussed on detailed unsteady aerodynamic inlet guide vanes (IGV)-Rotor interactions in a single stage transonic fan. In it's first part, the effects of boundary layer ingested inlet distortion on the unsteady flowfield are analyzed on a periodic quarter annulus domain of the IGV-Rotor fan stage, as they result in significant changes in the massflow rate, total pressure ratio and stage isentropic efficiency. The second part assessed the impact of axial IGV-Rotor spacing on the shock wave-boundary layer interaction. It dealt with the unsteady analysis of IGV-Rotor interactions in a single blade passage first stage transo (open full item for complete abstract)

Committee: Mark Turner Sc.D. (Committee Chair); Paul Orkwis Ph.D. (Committee Member); Prashant Khare Ph.D. (Committee Member); Shaaban Abdallah Ph.D. (Committee Member) Subjects: Aerospace Engineering

Master of Science, The Ohio State University, 2022, Aerospace Engineering

The effect of varying submergence on the transonic flow past canonical wall-mounted hemispheres is investigated at a freestream Mach number, M = 0.8 using Delayed Detached Eddy Simulations (DDES). Four submergence levels are considered ranging from a full hemisphere (100% exposed) to a highly submerged case where the equator is well below the waterline (40% exposed). Analysis of the mean characteristics indicates a reduction in strength and extent of dominant flow topology, including the horseshoe vortex and the counter-rotating vortices in the wake. Additionally, it is found that the mean line of surface separation moves downstream with submergence. This finding is validated with analysis of the unsteady streamwise shock position, which indicates the mean shock foot position also moved downstream with submergence. However, the frequency associated with the streamwise motion of the shock remains consistent between the cases in terms of a suitably non- dimensionalized Strouhal number, StD ∼ 0.26. The unsteady surface forces and modal analysis are used to quantify the effect of submergence on two correlated shock-wake modes, as these are dominant unsteady features in flow over full hemispheres. The first mode is the “breathing” mode, corresponding to spanwise symmetric wake shedding and correlates to streamwise shock oscillations. The second mode is the “shifting”, corresponding to spanwise anti-symmetric wake shedding and correlates with the spanwise rocking of the shock. Proper Orthogonal Decomposition (POD) is used to isolate and rank the different modes; as the hemisphere is submerged, there is an evident change in prominence from the anti-symmetric shifting to the symmetric breathing mode. Dynamic Mode Decomposition (DMD) is used to investigate the spectral content of these modes; the results show that the breathing mode collapses at a frequency of StD ∼ 0.26 for all cases, while the shifting mode extends over a broad frequency range of StD between 0.13 and 0.21.

Committee: Jack McNamara (Committee Member); Datta Gaitonde (Advisor) Subjects: Aerospace Engineering

Doctor of Philosophy, The Ohio State University, 2020, Mechanical Engineering

Gas turbine is an integrated part of modern aviation and power generation industry. The thermal efficiency of a gas turbine strongly depends on the turbine inlet temperature (TIT), and the turbine designers are continuously pushing the TIT to a higher value. Due to the increased freedom in additive manufacturing, the complex internal and external geometries of the turbine blade can be leveraged to utilize innovative cooling designs to address some of the shortcomings of current cooling technologies. The sweeping jet film cooling has shown some promise to be an effective method of cooling where the coolant can be brought very close to the blade surface due to its sweeping nature. A series of experiments were performed using a row of fluidic oscillators on a flat plate. Adiabatic cooling effectiveness, convective heat transfer coefficient, thermal field, and discharge coefficient were measured over a range of blowing ratios and freestream turbulence. Results were compared with a conventional shaped hole (777-hole), and the sweeping jet hole shows improved cooling performance in the lateral direction. Numerical simulation also confirmed that the sweeping jet creates two alternating vortices that do not have mutual interaction in time. When the jet sweeps to one side of the hole exit, it acts as a vortex generator as it interacts with the mainstream ow. This prevents the formation of the counter-rotating vortex pair (CRVP) and allows the coolant to spread in the lateral direction. The results obtained from the low speed at plate tests were utilized to design the sweeping jet film cooling hole for more representative turbine vane geometry. Experiments were performed in a low-speed linear cascade facility. Results showed that the sweeping jet hole has higher cooling effectiveness in the near hole region compared to the shaped hole at high blowing ratios. Next, a detailed experimental investigation of sweeping jet film cooling on the suction surface of a near engine scale (open full item for complete abstract)

Committee: Jeffrey Bons (Advisor); James Gregory (Committee Member); Randall Mathison (Committee Member); Ali Ameri (Committee Member) Subjects: Aerospace Engineering; Aerospace Materials; Engineering; Mechanical Engineering

Doctor of Philosophy, The Ohio State University, 2019, Aero/Astro Engineering

The purpose of the current work is to study the transient behavior of axial flow turbomachines in transonic environments. The first part is focused on the aerodynamic behavior of a coupled boundary layer ingesting inlet – distortion tolerant fan for use in next generation aircraft. The second part is focused on stall detection in transonic compressors. The reason these two are covered is that the fan and compressor of this future system would be mechanically linked. The fan behaves dynamically in the presence of a distortion and exhibits non-uniform flow downstream of the fan. An engine core, comprised of compressor stages, would have to operate in that distorted flow possibly inducing stall. Therefore, understanding the behavior of the fan and developing advanced stall detection methods are important for the successful implementation of these future systems. Part I – Aerodynamic Response of a Distortion Tolerant Fan Coupled to a Boundary Layer Ingesting Inlet Future aircraft designs are aimed at three main targets for more sustainable flight practices decreasing noise, emissions, and fuel burn. Boundary layer ingestion (BLI), as a part of aircraft's propulsion, is an innovative means of achieving improvements to all three targets. The BLI design incorporates engines integrated into the body of an aircraft. The engine ingests low momentum boundary layer flow that develops over the aircraft's surface when in operation. The advantages of such a system are reductions to weight, drag, noise, and increased propulsive efficiency over standard aircraft in service today. Experimental tests representing a blended wing body propulsor utilizing BLI were performed in NASA's 8ft by 6ft wind tunnel. These tests were aimed at obtaining the physically realizable benefits achievable for a boundary layer ingesting propulsor. The current work represents an effort to compare a simulated coupled boundary layer ingesting inlet and distortion tolerant fan with the experimental measure (open full item for complete abstract)

Committee: Jen-Ping Chen PhD (Advisor); Milind Bakhle PhD (Committee Member); Han-Wei Shen PhD (Committee Member); Datta Gaitonde PhD (Committee Member); Michael Dunn PhD (Committee Member) Subjects: Aerospace Engineering

MS, University of Cincinnati, 2017, Engineering and Applied Science: Aerospace Engineering

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.

Committee: Paul Orkwis Ph.D. (Committee Chair); Shaaban Abdallah Ph.D. (Committee Member); Mark Turner Sc.D. (Committee Member) Subjects: Aerospace Materials

Doctor of Philosophy, The Ohio State University, 1984, Graduate School

Committee: Not Provided (Other) Subjects: Engineering

Doctor of Philosophy, The Ohio State University, 1982, Graduate School

Committee: Not Provided (Other) Subjects: Engineering

Master of Science in Mechanical Engineering (MSME), Wright State University, 2016, Mechanical Engineering

A computational study was conducted to understand the influence of aircraft inlet distortion flow on the unsteady aerodynamic loading of a gas turbine fan stage. A single stage, transonic fan design with no inlet guide vanes was modeled with a commercial, computational fluid dynamics solver, STAR-CCM+, using the harmonic balance technique. The baseline inlet boundary condition applied to the model is consistent with that of a homeomorphic variant of the M2129 S-duct, and exhibited stagnation pressure distortion and a swirl pattern. The baseline inlet flow was decomposed and parameterized into a set of inlet boundary conditions which were individually applied in a series of computational runs. The parametric effect of the swirl velocity on the unsteady aerodynamic loading of the fan blade was studied. The flow structure at 90% span was investigated and revealed localized flow patterns in the blade row. A Fourier analysis revealed that the inlet distortion did not simply convect with the flow, but changed strength along the streamline, redistributing the spectral energy of the pressure and velocity components, and was dependent on the swirl parameter value. The inlet distortion also caused the passage shock to move up to 12% of the blade chord. Finally, the first six harmonics of the inlet distortion contributed significantly to the aerodynamic loading of the transonic fan.

Committee: Mitch Wolff Ph.D. (Advisor); David Johnston Ph.D. (Committee Member); George Huang Ph.D. (Committee Member); Joseph Slater Ph.D. (Other) Subjects: Aerospace 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

Master of Science in Engineering (MSEgr), Wright State University, 2003, Mechanical Engineering

Hutton, Timothy. M.S. Egr., Department of Mechanical Engineering, Wright State University, 2003. Innovative Forced Response Analysis Method Applied To A Transonic Compressor. A set of inlet guide vane (IGV) unsteady surface pressure measurements of a transonic compressor is presented. Using a flexible pressure sensor array, unsteady IGV suction-surface and pressure-surface pressures are acquired for six spanwise by five chordwise locations for various speed lines and throttle settings. Measurements from this sensor array are used to investigate unsteady vane/blade interaction aeromechanical forcing functions in a modern, highly loaded compressor stage. A significant effect is shown on the unsteady forced response of the IGV with changes in compressor operating point and IGV/rotor axial spacing for various span and chord locations. In particular, variations in the compressor operating point (i.e., mass flow rate and pressure ratio) cause change in both the magnitude and phase of the forced response, with the near-stall operating point producing the highest response. Changes in the axial spacing between the IGV and rotor rows from 12% to 26% of the IGV chord resulted in a 50% reduction in the magnitude of the forced response. A significant variation in the forced response with span is noted, especially at the 5% span location where the rotor relative flow is subsonic. In this region, changes in the operating point and axial spacing had a negligible effect on the forced response of the IGV. An innovative data reduction/analysis method is presented to quantify and statistically analyze the degree of blade-to-blade variations in the measured aerodynamic forcing functions obtained by turbomachinery experimentation. This method is used to analyze experimental data of IGV surface unsteady pressure response due to the aerodynamic forcing function produced by the downstream transonic compressor rotor with (1) factory-whole blades and (2) trimmed (blended) blades resulting fro (open full item for complete abstract)

Committee: J. Wolff (Advisor) Subjects: Engineering, Mechanical

MS, University of Cincinnati, 2007, Engineering : Aerospace Engineering

This thesis deals with the effects of the two types of sidewalls boundary conditions, namely slip and periodic, impact the computed acoustic field of transonic cavity. A Menter-SST based detached eddy simulation (DES) is adopted as the main turbulence modeling methodology. The hybrid DES approach for modeling turbulence enables the achievement of the desired eddy resolution in the massively separated flow region within the cavity without resorting to the excessive grid requirements of large eddy simulations (LES). Furthermore, the numerical resolution of the wide dynamic range involved is accomplished through the use of Roe third order scheme. Near-wall simulation is approximated by a wall function to reduce computational costs for the unsteady high Reynolds number flow. Simulations were performed on a multi-zone structured domain by parallel processing using the WIND code. Computational results are presented for the instantaneous velocity, vorticity, Mach number, pressure fluctuations and for the sound pressure level (SPL) and turbulent kinetic energy (TKE) spectra within the cavity. The spectra, which were generated using the maximum entropy method, are presented for the slip and periodic sidewall boundary condition cases. A comparison with available experimental results, indicate that the slip sidewall boundary conditions produce SPL spectra in better agreement with the experimental results. A summary of computational costs is presented to demonstrate the savings compared to prior low Reynolds number results obtained using DES and LES methods. Time series analyses were also taken on sidewall points and compared of the two cases with different boundary conditions.

Committee: Dr. Awatef Hamed (Advisor) Subjects: Engineering, Aerospace

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

Experimental and numerical investigations have shown that flow in the tip region of high-speed, low aspect ratio, transonic compressors is detrimental to aerodynamic efficiency and operating range. Various computational fluid dynamic (CFD) models were compared to determine their influence on off-design predictions and tip flow structures. The models compared were: 1) steady isolated rotor, 2) unsteady isolated rotor, 3) unsteady stage. The κ-ε and κ-ω turbulence models were also compared. The near stall predictions were oscillatory in all cases. The largest variation in numerical results occurred due to choice of turbulence model. The inability to obtain a converged steady-state solution at low flow rates indicates the significance of an unsteady phenomenon present in the flow field indicating the need for unsteady flow simulations at off-design conditions. The current investigation has shown that unsteady isolated rotor simulations provide the most practical balance of accuracy in prediction and computational cost.

Committee: Dr. Paul D. Orkwis (Advisor) Subjects:

Doctor of Philosophy, The Ohio State University, 2009, Aeronautical and Astronautical Engineering

This is the first 3-D unsteady RANS simulation of a highly loaded transonicturbine stage and results are compared to steady calculations and experiments. A low Reynolds number k-ε turbulence model is employed to provide closure for the RANS system. Phase-lag is used in the tangential direction to account for stator-rotor interaction. Due to the highly loaded characteristics of the stage, inviscid effects dominate the flowfield downstream of the rotor leading edge minimizing the effect of segregation to the leading edge region of the rotor blade. Unsteadiness was observed at the tip surface that results in intermittent 'hot spots'. It is demonstrated that unsteadiness in the tip gap is governed by both inviscid and viscous effects due to shock-boundary layer interaction and is not heavily dependent on pressure ratio across the tip gap. This is contrary to published observations that have primarily dealt with subsonic tip flows. The high relative Mach numbers in the tip gap lead to a choking of the leakage flow that translates to a relative attenuation of losses at higher loading. The efficacy of a new tip geometry is discussed to minimize heat flux at the tip while maintaining choked conditions. Simulated heat flux and pressure on the blade and hub agree favorably with experiment and literature. The time-averaged simulation provides a more conservative estimate of heat flux than the steady simulation. The shock structure formed due to stator-rotor interaction is analyzed. A preprocessor has also been developed as a conduit between the unstructured multi-block grid generation software GridPro and the CFD code TURBO.

Committee: Jen-Ping Chen PhD (Advisor); Ali Ameri PhD (Committee Member); Meyer Benzakein PhD (Committee Member); Jeffrey Bons PhD (Committee Member); Henry Busby PhD (Committee Member) Subjects: Engineering; Physics

MS, University of Cincinnati, 2012, Engineering and Applied Science: Aerospace Engineering

The primary objective of this thesis is to advance Proper Orthogonal Decomposition (POD) methods to quantify similarities and differences between the turbulent structures in the mixing layer of transonic jets issued through baseline axisymmetric conical nozzle and ones with chevrons. The analysis is done using flow velocity data obtained through Particle Image Velocimetry (PIV). Two chevron nozzles with penetration levels at 2% and 4% are used. The mean velocity and TKE results show that chevrons reduce the potential core length, increase the jet spread, increase TKE levels immediately after the nozzle exit, and reduce TKE levels downstream. POD is used to further quantify these results. First 35 converged POD modes are investigated since these modes comprise of the majority of the large scale structures. The similarities and differences between the POD modes from different data sets vary along the jet. The study of the POD modes along the jet showed the growth of the structures in both size and strength in the downstream direction. The growth rates of the modes are suppressed by the chevrons and high penetration chevrons are more effective. This trend is also noticed in the energy distribution study among the modes. Despite having both the axial and radial modes energized by the chevrons near the nozzle, the stabilization of the modes resulted in lower energy contents in the downstream regions. Projection of the PIV images from the chevron configurations onto the baseline POD modes showed that both near the nozzle and downstream the highest energy containing baseline modes are attenuated by the chevrons, but the radial modes in the near-nozzle region. A metric used in order to quantify the similarities among the mode shapes between different flow data sets showed that in the near nozzle region the correlation between the modes corresponding to the chevron nozzles and baseline nozzle is very low due to the increased mixing. Correlation increases as the flow (open full item for complete abstract)

Committee: Ephraim Gutmark PhD DSc (Committee Chair); Kelly Cohen PhD (Committee Member); Jeffrey Kastner PhD (Committee Member) Subjects: Aerospace Materials