Doctor of Philosophy (Ph.D.), University of Dayton, 2024, Engineering

With the increasing interest in electric vertical takeoff and landing air vehicles and small-scale Unmanned Air Vehicles, many novel design concepts favor the fixed-pitch-propeller as the primary propulsion system due to its simplicity and reliability. This expands the application scenario of the fixed-pitch propeller from axial forward flight to edgewise flight conditions. The current study investigated the changes in its performance when operating at higher incidence angle conditions as well as the proximity effects of the propellers in these conditions. It is hypothesized that the propeller performance under various conditions and proximities can be reasonably predicted by modeling the changes in the inflow angle of the propeller. This hypothesis was tested using three major steps. First, a relationship between inflow angle, propeller inclination angle, and advance ratio was established using a series of experimental investigations. Second, this relationship was used to predict the performance of two propellers in tandem configuration with various horizontal and vertical offset distances. Third, the same model was used to predict the ceiling effect of the propeller at different incidence angles and advance ratios. All experiments were conducted at the University of Dayton Low-Speed Wind Tunnel (UD-LSWT) Laboratory under its open jet configuration. Force-based experiments, flow visualization as well as phase-locked Particle Image Velocimetry (PIV) experiments were conducted for all investigations. The changes in propeller performance at various flight conditions were quantified and several normalization methods were successfully employed indicating the predictability of various propeller forces and moments. A novel propeller axial thrust prediction model was proposed considering the propeller performance as a summation of propeller-like components and wing-like component, with an overall error of less than 8.3%. Flow visualization and PIV results confirmed the (open full item for complete abstract)

Committee: Sidaard Gunasekaran (Committee Chair); Michael OL (Committee Member); Markus Rumpfkeil (Committee Member); Aaron Altman (Committee Member) Subjects: Aerospace Engineering

Master of Science (M.S.), University of Dayton, 2024, Aerospace Engineering

Recent advancements in battery technology have led to an increase in the development of electric Vertical Takeoff and Landing (eVTOL) vehicles, typically using electrically-powered propellers to generate both lift and thrust. These vehicles typically operate in low-altitude, and limited-space conditions in urban environments. Unsteady flows from building wakes or atmospheric boundary layer effects raise concern to the stability of eVTOL-capable aircraft under normal operating conditions and during transition from vertical to forward flight and vice-versa in population dense areas. Although all types of unsteady flows have been studied for decades, little has been published on the influence of unsteady flow on a propeller-wing system. Understanding of this system is crucial to ensuring the safety of not only the passengers of these VTOL aircraft, but also the safety of the public. Investigation into powered wing response to streamwise gust encounters was conducted through various propeller locations, angles of attack, reduced frequencies, and thrust levels. All experiments were run at the University of Dayton Low Speed Wind Tunnel (UD-LSWT) in its open-jet configuration. The shuttering system downstream of the test section consists of a set of rotating louvers that change angle to effectively change the blockage ratio of the wind tunnel. Different louver angles and actuation frequencies provide different velocities and reduced frequencies. Particle Image Velocimetry (PIV) was conducted on the freestream flow during actuation of the louvers to spatially characterize the angle of attack variation throughout the test section. The results from PIV were used to determine the optimal testing location and wing size for the test article. The wing was designed to be modular, accepting a number of different propeller Distribution Statement A: Approved for Public Release; Distribution is Unlimited. PA# AFRL-2024-2083 4 locations. Four total configurations were considered – (open full item for complete abstract)

Committee: Sidaard Gunasekaran (Committee Chair); Michael Mongin (Committee Member); Albert Medina (Committee Member); Markus Rumpfkeil (Committee Member) Subjects: Aerospace Engineering

Master of Science, The Ohio State University, 2024, Aerospace Engineering

Uncrewed aerial vehicles (UAVs) have seen an explosion in development over the last decade. Advances in computing and communication technologies have allowed for rapid innovation in remote and autonomous flight controls, leading to development of new aerospace platforms. Without humans on board, vehicles have restrictions removed that can lead to unique configurations; for example, UAVs can be much smaller and more maneuverable than other platforms. An uncrewed platform is not limited by the maximum G-force its pilot can sustain. Relatively low costs further provide opportunities for innovation in UAV design. Thanks to reduced costs per vehicle and autonomous control systems, UAVs could operate with smaller error margins near the regimes of failure. Swept wings are selected in aircraft design for their aerodynamic performance at high Mach number. By reducing the component of freestream velocity incident on the wing in the direction normal to the leading edge, wing sweep increases the critical Mach number where a marked increase in drag is observed. While wing sweep is desirable for flight at high Mach numbers, a swept wing geometry can pose aerodynamic challenges in low speed, high angle of attack flight. The swept geometry develops a spanwise pressure gradient that is not observed on straight wings. The resulting spanwise flow develops a natural boundary layer separation control that is most effective the wing root. As a swept wing's angle of attack is increased, flow will initially separate near the wingtip and trailing edge. As the angle of attack continues to increase, separation will propagate toward the wing root and leading edge. Taking advantage of these fluid mechanics, a flow control system that interrupts spanwise flow over a swept wing's suction surface can improve aerodynamic performance at high angles of attack. Because the interruption is not needed for high-speed flight, an active flow control method is desirable. The unsteady aerodynamic eff (open full item for complete abstract)

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

Doctor of Philosophy, The Ohio State University, 1980, Graduate School

Committee: Not Provided (Other) Subjects: Engineering

Doctor of Philosophy, The Ohio State University, 2015, Aero/Astro Engineering

One of the most pervasive threats to aircraft controllability is wing stall, a condition associated with loss of lift due to separation of air flow from the wing surface at high angles of attack. A recognized need for improved upset recovery training in extended-envelope flight simulators is a physical understanding of the post-stall aerodynamic environment, particularly key flow phenomena which influence the vehicle trajectory. Large-scale flow structures known as stall cells, which scale with the wing chord and are spatially-periodic along the span, have been previously observed on post-stall airfoils with trailing-edge separation present. Despite extensive documentation of stall cells in the literature, the physical mechanisms behind their formation and evolution have proven to be elusive. The undertaken study has sought to characterize the inherently turbulent separated flow existing above the wing surface with cell formation present. In particular, the question of how the unsteady separated flow may interact with the wing to produce time-averaged cellular surface patterns is considered. Time-resolved, two-component particle image velocimetry measurements were acquired at the plane of symmetry of a single stall cell formed on an extruded NACA 0015 airfoil model at chord Reynolds number of 560,000 to obtain insight into the time-dependent flow structure. The evolution of flow unsteadiness was analyzed over a static angle-of-attack range covering the narrow post-stall regime in which stall cells have been observed. Spectral analysis of velocity fields acquired near the stall angle confirmed a low-frequency flow oscillation previously detected in pointwise surface measurements by Yon and Katz (1998), corresponding to a Strouhal number of 0.042 based on frontal projected chord height. Probability density functions of the streamwise velocity component were used to estimate the convective speed of this mode at approximately half the free-stream velocity, in (open full item for complete abstract)

Committee: James Gregory Ph.D. (Advisor); Jeffrey Bons Ph.D. (Committee Member); Mo Samimy Ph.D. (Committee Member); Jen-Ping Chen Ph.D. (Committee Member) Subjects: Aerospace Engineering; Fluid Dynamics

Master of Science in Mechanical Engineering, Cleveland State University, 2013, Fenn College of Engineering

The purpose of this thesis is to document the impact of incidence angle and Reynolds number variations on the 3-D flow field and midspan loss and turning of a 2-D section of a variable-speed power-turbine (VSPT) rotor blade. Aerodynamic measurements were obtained in a transonic linear cascade at NASA Glenn Research Center in Cleveland, OH. Steady-state data were obtained for ten incidence angles ranging from +15.8&00B0; to -51.0&00B0;. At each angle, data were acquired at five flow conditions with the exit Reynolds number (based on axial chord) varying over an order-of-magnitude from 2.12 &00D7; 10^5 to 2.12 &00D7; 10^6. Data were obtained at the design exit Mach number of 0.72 and at a reduced exit Mach number of 0.35 as required to achieve the lowest Reynolds number. Midspan total-pressure and exit flow angle data were acquired using a five-hole pitch/yaw probe surveyed on a plane located 7.0 percent axial-chord downstream of the blade trailing edge plane. The survey spanned three blade passages. Additionally, three-dimensional half-span flow fields were examined with additional probe survey data acquired at 26 span locations for two key incidence angles of +5.8&00B0; and -36.7&00B0;. Survey data near the endwall were acquired with a three-hole boundary-layer probe. The data were integrated to determine average exit total-pressure and flow angle as functions of incidence and flow conditions. The data set also includes blade static pressures measured on four spanwise planes and endwall static pressures.

Committee: Mounir Ibrahim PhD (Committee Chair); Miron Kaufman PhD (Committee Member); Ralph Volino PhD (Committee Member) Subjects: Aerospace Engineering; Mechanical Engineering

Master of Science (M.S.), University of Dayton, 2010, Aerospace Engineering

An investigation into wake and solid blockage effects of Vertical-Axis Wind Turbines (VAWTs) in closed test-section wind tunnel testing is described. Static wall pressures have been used to derive velocity increments along a wind tunnel test-section which in-turn are applied to provide evidence of wake interference characteristics of rotating bodies interacting within this spatially restricted domain. Vertical-axis wind turbines present a unique aerodynamic obstruction in wind tunnel testing whose blockage effects have not been extensively investigated.The flow-field surrounding these wind turbines is asymmetric, periodic, unsteady, separated and highly turbulent. Static pressure measurements are taken along a test-section sidewall to provide a pressure signature of the test models under varying rotor tip-speed ratios (freestream conditions and model RPM's). To provide some guidance on the scaling of the combined effects of wake and solid blockage, wake characteristics and VAWT performance produced by the same vertical-axis wind turbine concept have been tested at different physical scales in two different wind tunnels. This investigation provides evidence of the effects of large wall interactions and wake propagation caused by these models at well below generally accepted standard blockage figures.

Committee: Aaron Altman PhD (Committee Chair); Jewel Barlow PhD (Committee Member); Eric Lang PhD (Committee Member) Subjects: Aerospace Materials; Engineering; Experiments; Fluid Dynamics; Mechanical Engineering

Master of Science, The Ohio State University, 2006, Graduate School

Committee: Not Provided (Other) Subjects:

Master of Science, The Ohio State University, 1963, Graduate School

Committee: Not Provided (Other) Subjects:

Master of Science, The Ohio State University, 2007, Graduate School

Committee: Not Provided (Other) Subjects:

Master of Science, The Ohio State University, 1966, Graduate School

Committee: Not Provided (Other) Subjects:

Master of Science, The Ohio State University, 1971, Graduate School

Committee: Not Provided (Other) Subjects:

Master of Science, The Ohio State University, 1965, Graduate School

Committee: Not Provided (Other) Subjects:

Master of Science, The Ohio State University, 2007, Graduate School

Committee: Not Provided (Other) Subjects:

Master of Science, The Ohio State University, 1966, Graduate School

Committee: Not Provided (Other) Subjects:

Master of Science, The Ohio State University, 1965, Graduate School

Committee: Not Provided (Other) Subjects:

Master of Science, The Ohio State University, 1964, Graduate School

Committee: Not Provided (Other) Subjects:

Master of Science (M.S.), University of Dayton, 2024, Aerospace Engineering

Traditional conceptual-level aerodynamic analysis is limited to empirical and/or inviscid models due to considerations of computational cost and complexity. There is a distinct desire to incorporate higher-fidelity analysis into the conceptual-design process as early as possible. This work seeks to enable the use of high-fidelity data by developing and applying multi-fidelity surrogate models that can efficiently predict the underlying response of a system with high accuracy. To that end, a novel form of the multi-fidelity polynomial chaos expansion (PCE) method is introduced, extending the surrogate modeling technique to accept three distinct fidelities of input. The PCE implementation is evaluated for a series of analytical test functions, showing excellent accuracy in creating multi-fidelity surrogate models. Aerodynamic analysis of a generic hypersonic vehicle (GHV) is performed using three codes of increasing fidelity: CBAERO (panel code), Cart3D (Euler), and FUN3D (RANS). The multi-fidelity PCE technique is used to model the aerodynamic responses of the GHV over a broad, five-dimensional input domain defined by Mach number, dynamic pressure, angle of attack, and left and right control surface settings. Mono-, bi-, and tri-fidelity PCE surrogates are generated and evaluated against a high-fidelity “truth” database to assess the global error of the surrogates focusing on the prediction of lift, drag, and pitching moment coefficients. Both monofidelity and multi-fidelity surrogates show excellent predictive capabilities. Multi-fidelity PCE models show significant promise, generating aerodynamic databases anchored to RANS fidelity at a fraction of the cost of direct evaluation.

Committee: Markus Rumpfkeil (Advisor); Jose Camberos (Committee Member); Timothy Eymann (Committee Member) Subjects: Aerospace Engineering

Master of Science, The Ohio State University, 2024, Aerospace Engineering

Motivated by the ongoing development of Turbine Based Combined Cycle systems, special interest was given to the low-speed characteristics of a vehicle designed for hypersonic operation. More specifically, the subsonic performance of the low-speed inlet for a TBCC vehicle was the primary focus of this study due to its integral role in take-off and landing operation. A wind tunnel model of a hypersonic scramjet inlet previously tested at hypersonic speeds was characterized for low-speed operation in a subsonic 3'x5' wind tunnel. This model contained a variable thrust simulator powered by high pressure air and was tested over a range of 0-20-degrees angle of attack at two Reynolds numbers, 250k and 500k. Static pressure data was collected along the model's centerline. The resulting data set was used to assess the validity of a computational fluid dynamic model within ANSYS Fluent setup to match wind tunnel conditions. The model was not able to accurately predict the experimental results due to oversimplifications and assumptions made. Despite this shortcoming, the model was still used to analyze the performance of an inlet optimized for Mach 4 that was integrated into the existing geometry of the wind tunnel model. This analysis emphasized the importance of subsonic inlet design and provides the framework for future development and design iteration.

Committee: Clifford Whitfield (Advisor); Matthew McCrink (Committee Member) Subjects: Aerospace Engineering

Master of Science, The Ohio State University, 2024, Aerospace Engineering

Thrust vectoring (TV) is the ability to manipulate the directivity of the primary jet to provide a cross-stream force off the primary jet axis. TV can enable desirable flight regimes such as hyper-maneuverability and short/vertical take-off and landing. Modern conventional TV methods utilize a physical mechanism to mechanically deflect the jet flow and change the thrust direction. This method is both heavy and mechanically complex, especially for an axisymmetric jet. A novel approach to TV is explored in this paper by investigating localized arc filament plasma actuators' (LAFPAs) ability to impart a TV force on a subsonic, axisymmetric jet by attaching the flow to a radially expanding surface (termed “reaction surface”), at the jet exit. LAFPAs will be used to asymmetrically control the entrainment of the jet to provide a deflection of the jet via the conservation of momentum. The deflected jet will then attach to the reaction surface via the Coanda effect. The jet flow was interrogated at baseline and excited cases with a static pressure array located 0.75 jet diameters downstream of the actuators at 70% of the chord of the reaction surface and with cross-stream particle image velocimetry (PIV) located 3 jet diameters downstream of the actuators. Two jet Mach (Mj) values were assessed, Mj = 0.48 and 0.9. The results show that the LAFPAs create a repeatable and significant asymmetric pressure profile trend with respect to excitation frequency. In general, low excitation frequencies provide an asymmetric azimuthal pressure profile that corresponds to a vectored thrust force towards the active actuators, while high excitation frequencies provide an asymmetric azimuthal pressure profile that corresponds to a vectored thrust force away from the active actuators. Cross-stream PIV flow field measurements show that the asymmetry in the azimuthal pressure profile is not as significant as would be desirable for thrust-vectoring applications. However, the PIV results do show (open full item for complete abstract)

Committee: Dr. Nathan Webb (Advisor); Dr. Mo Samimy (Committee Member) Subjects: Aerospace Engineering