Theses and Dissertations from UMD
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New submissions to the thesis/dissertation collections are added automatically as they are received from the Graduate School. Currently, the Graduate School deposits all theses and dissertations from a given semester after the official graduation date. This means that there may be up to a 4 month delay in the appearance of a give thesis/dissertation in DRUM
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Item AEROSOL EFFECTS IN HIGH SUPERSONIC FLOWS(2024) Schoneich, Antonio Giovanni; Laurence, Stuart J; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)The understanding of high-speed aerodynamics is becoming evermore pertinent with thegrowth of space tourism, continued interest in space exploration, and pursuit of advanced highspeed aircraft for both military and commercial use. For initial investigations, ground test facilities are preferred to flight tests as they are far cheaper and carry significantly less risk, although wind tunnels can only replicate a subset of the conditions experienced in actual flight. One of these conditions that has not been adequately captured in wind tunnels is the effect of particulates in the atmosphere. Typical wind tunnels use a pure, clean gas (air, nitrogen, etc.) for testing, but this does notcapture the aerosolized nature of the atmosphere, where humidity and condensation can produce a distribution of liquid droplet sizes ranging from the average rain drop of 2mm to sub-micron diameter particles. Similarly, volcanic eruptions and ever-present wildfires result in solid particles exhibiting a variety of species and sizes that are transported to every layer of the atmosphere. At supersonic speeds, encounters with particulates have been shown to lead to detrimental effects, such as material erosion and boundary layer transition. Previous attempts to study this problem in wind tunnels have focused mainly on sub-micronsized solid particles, since aerosol settling time is a major limiting factor. On the other hand, most high-speed experiments involving large liquid droplet impacts have been carried out in gas guns or ballistic ranges due to the difficulty of trying to accelerate a droplet to high speeds without causing it to break up. While these facilities can be used to study impacts, the moving model means that detailed aerodynamic studies are nearly impossible, leading to a large gap in knowledge. To perform high-speed wind tunnel testing with liquid aerosols representative of cloud-likeenvironments (5-20 μm), a Mach-4 facility, referred to as the Multi-phase Investigations Supersonic Tunnel (MIST) has been designed and developed at the University of Maryland (capable of producing supersonic, particle-laden flows). This range of aerosol sizes makes MIST a unique facility with significant potential for expanding the state of the art in high-speed multi-phase flows. The present work discusses the design and characterization of MIST as well as two major experimental investigations carried out using this new facility. The first investigation examines the force augmentation on a free-flying sphere exposed to supersonic, particle-laden flows. Freeflight measurements are performed with five different particle size and concentration combinations. When comparing the results for particle-free flow in the same facility, the drag coefficient of the sphere was shown to be 1.75-4.5% greater for all multi-phase cases; this is significantly higher than simple estimates based on the increased momentum flux in the freestream would indicate. In addition to force measurements, an experimental investigation into the effect of particle-ladenflows on boundary-layer transition was conducted. It is important to characterize the disturbance environment in wind tunnels since they typically do not represent the levels in atmospheric flight and can lead to earlier onset of boundary-layer transition. In performing such measurements using a single-point Focused Laser Differential Interferometer, it was discovered that the presence of particles in the flow could significantly attenuate the acoustic disturbances generated by the wind tunnel. This finding was further reinforced when investigating the boundary-layer transition on a 5◦ half-angle, sharp cone using high-speed schlieren visualization. For each case presented in this work, the boundary-layer disturbance amplitudes were reduced and transition Reynolds numbers increased in the particle-laden flow cases. This was contrary to expectations, given that prior numerical studies have indicated that particles can induce early transition. These findings potentially open a path to substantially reduce freestream disturbance levels in conventional hypersonic wind tunnels.Item INVESTIGATION OF COMPOUND ROTORCRAFT AEROMECHANICS THROUGH WIND-TUNNEL TESTING AND ANALYSIS(2022) Maurya, Shashank; Datta, Anubhav; Chopra, Inderjit; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)The aeromechanics of a slowed-rotor compound rotorcraft is investigated through wind-tunnel testing and comprehensive analysis. The emphasis is on a lift-offset wing compound with a hingeless rotor configuration. A new Maryland Compound Rig is developed and instrumented for wind-tunnel testing and an in-house rotor comprehensive code is modified and expanded for compound rotorcraft analysis. The compound rig consists of a lift compound model and a propeller model. The lift compound model consists of an interchangeable hub (articulated or hingeless), a fuselage, a half-wing of 70% rotor radius on the retreating side. The wing has a dedicated load cell and multiple attachment points relative to the rotor hub (16%R, 24%R, and 32%R and 5%R aft of the hub). The rotor diameter is 5.7-ft. The rotor has four blades with NACA 0012 airfoils with no twist and no taper. The wing incidence angle is variable between 0 to 12 degrees. The wing has a linearly varying thickness with symmetric airfoils NACA 0015 at the tip and NACA 0020 at the root. Sensors can measure rotor hub forces and moments, wing root forces and moments, blade pitch angles, structural loads (flap bending moment, lagbending moment, and torsional moment) at 25%R, pitch link loads, and hub vibratory loads. Wind tunnel tests are conducted up to advance ratio 0.7 for lift compound with half-wing at wing incidence angles of 4 and 8 degrees and compared with an isolated rotor. Hover tests are conducted up to tip Mach number of 0.5 to measure download penalty with the wing at various positions. The University of Maryland Advanced Rotorcraft Code (UMARC) is modified for compound rotorcraft analysis code. Aerodynamic models for the wing and the propeller are integrated. A recently developed Maryland Free Wake model is integrated, which can model the wake interaction between unequal and inharmonic speed rotor, wing, and propeller. The analysis is then validated with the test data. The validated analysis is used to analyze the US Army hypothetical full-scale aircraft. The compound rotorcraft is categorized into multiple configurations in a systematic manner to find the extreme limits of speed and efficiency of each. The key conclusions are: 1) slowing the rotor or compounding the configuration provide no benefit individually; they must be accomplished together, 2) Half-Wing is more beneficial if a lift-offset hingeless rotor is used, 3) hover download penalty is only 3% of net thrust, and this penalty can be predicted satisfactorily by free wake, 4) the main rotor wake interaction is more pronounced on the wing and less on the propeller, 5) the validated analysis indicates a speed of 240 knots may be possible with 20% RPM reduction along with a wing and propeller, if structural weights allow, and 6) the oscillatory and vibratory lag moments and in-plane hub loads may be significantly reduced by compounding.Item SURROGATE MODELING AND CHARACTERIZATION OF BLADE-WAKE INTERACTION NOISE FOR HOVERING SUAS ROTORS USING ARTIFICIAL NEURAL NETWORKS(2022) Thurman, Christopher; Baeder, James; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)This work illustrates the use of artificial neural network modeling to study and characterize broadband blade-wake interaction noise from hovering sUAS rotors subject to varying airfoil geometries, rotor geometries, and operating conditions. Design of Experiments was used to create input feature spaces over 9 input features: the number of rotor blades, rotor size, rotor speed, the amount of blade twist, blade taper ratio, tip chord length, collective pitch, airfoil camber, and airfoil thickness. A high-fidelity strategy was then implemented at the discrete data points defined by the designed input feature spaces to design airfoils and rotor blades, predict the unsteady rotor aerodynamics and aeroacoustics, and isolate the blade-wake interaction noise from the acoustic broadband noise, which was then used for prediction model training and validation. An artificial neural network tool was developed and implemented into NASA's ANOPP2 code and was used to identify an optimal prediction model for the nonlinear functional relationship between the 9 input features and blade-wake interaction noise. This optimal artificial neural network was then validated over test data, and exhibited prediction accuracy over 91% for data previously unseen by the model. First- and second-order sensitivity analyses were then conducted using the developed artificial neural network tool and it was seen that input features which serve to directly modify the thrust coefficient, such as airfoil camber and collective pitch, had a dominant effect over blade-wake interaction noise, followed by second-order interaction effects related to the mean rotor solidity. The optimal prediction model along with aerodynamic simulations were used to further study the effect of varying input features on blade-wake interaction noise and three types of blade-wake interaction noise were identified. Blade-wake interaction noise caused by impingement of the turbulence entrained in a tip vortex on the leading edge of a subsequent rotor blade showed to be the most prominent type of blade-wake interaction noise, exhibiting an acoustic contribution upward of 7 dB. Blade-wake interaction noise caused by a direct impingement of a tip vortex on the leading edge of a subsequent rotor blade had the second largest acoustic significance, exhibiting roughly 6 dB of broadband noise. The third, and least significant type of blade-wake interaction noise was shown to be caused by impingement of a blade-wake sheet on the mid-span of a subsequent rotor. This last type of blade-wake interaction noise was seen to only occur in the turbulent-wake operating state and possibly mild vertical descent conditions, and had approximately a 2.5 dB acoustic contribution.Item Aeroelastic Stability Analysis of a Wing with a Variable Cant Angle Winglet(2020) Mondragon Gomez, Jose Mauricio; Hubbard, James E; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)Currently, multiple air vehicles employ wing shape change to enhance their performance and achieve mission adaptability in different environments inside the Earth's atmosphere. This concept has been around since the dawn of aviation. In 1903, the Wright brothers implemented wing warping to control their aircraft during flight. Subsequently, a variety of techniques and devices have used to achieve wing shape change and make the vehicles more versatile. For example, they include variable wing sweep, folding wing tips, and variable camber. However, aeroelasticity has played in important role in these developments. Thus, this work focuses on the aeroelastic analysis and understanding of the fundamental physics of the flutter mechanism of a wing equipped with a variable cant angle winglet. Two methods are applied to model the wingletted wing system. The Rayleigh Ritz method is the first technique used to model the system. This method involves the implementation of a shape function to represent the entire structure. The second method used in the analysis is the Finite Element Analysis. In this formulation, the wing structure is divided into elements and elemental functions are used for local interpolation. Strip theory is used to model the spanwise aerodynamic loading. In addition, steady, quasi-steady, and unsteady aerodynamic models are used, each with different levels of complexity. Both the structural and aerodynamic models were coupled to generate four dynamic aeroelastic equations that represent the continuous system. Those equations were used to model the system and perform a dynamic aeroelastic analysis. The results indicate that having a vehicle equipped with a variable cant angle winglet can be favorable. It can increase flutter speed and expand its flight envelope. Moreover, when the winglet length is greater than 50% the length of the wing section and the cant angle greater than 50 degrees, the second torsional mode of vibration becomes unstable. Whereas, the first mode remains marginally stable. Thus, the second mode has become the critical mode that leads to structural failure. In this case, that phenomenon is referred as mode switching.Item Wind Tunnel Test on Slowed Rotor Aeroechanics at High Advance Ratios(2020) Wang, Xing; Chopra, Inderjit; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)In forward flight, slowing down a rotor alleviates compressibility effects on the advancing side, extending the cruise speed limit and inducing high advance ratio flight regime. To investigate the aerodynamic phenomena at high advance ratios and provide data for the validation of analysis tools, a series of wind tunnel tests were conducted progressively with a 33.5-in radius, 4-bladed Mach-scaled rotor in the Glenn L. Martin Wind Tunnel. In the first stage of the research, a wind tunnel test was carried out at high advance ratios with highly similar, non-instrumented blades and on-hub control angle measurements, in order to gain a baseline performance and control dataset with minimum error due to blade structural dissimilarity and pitch angle discrepancy. The tests were conducted at advance ratios of 0.3 to 0.9, and a parametric study on shaft tilt was conducted at $0^\circ$ and $\pm 4^\circ$ shaft tilt angles. The test data were then compared with those of previous tests and with the predictions of the in-house comprehensive analysis UMARC. The airload results were investigated using comprehensive analysis to gain insights on the influences of advance ratio and shaft tilt angle on rotor performance and hub vibratory loads. Results indicate that the thrust benefit from backward shaft tilt is dependent on the change in the inflow condition and the induced angle of attack increment, and the reverse flow region at high advance ratios is the major contributor to changes in shaft torque and horizontal force. In the second stage of the research, the rotor blades were instrumented with pressure sensors and strain gauges at 30\% radius, and pressure data were acquired to calculate the sectional airloads by surface integration up to an advance ratio of 0.8. The test results of blade airloads and structural loads were compared with the predictions of comprehensive analysis (UMARC and PrasadUM) and CFD/CSD coupled analysis (PrasadUM/HAMSTR). The focus was on the data correlation between experimental pressure, airload and structural load data and the CFD/CSD predicted results at various collective and shaft tilt settings. Overall, the data correlation was found satisfactory, and the study provided some insights into the aerodynamic mechanisms that affect the rotor airload and performance, in particular the mechanisms of backward shaft tilt, hub/shaft wake and the formation of dynamic stall in the reverse flow region. The next stage focused on hingeless rotor with lift offset. Previous wind tunnel tests have shown that an articulated rotor trimmed to zero hub moment generates limited thrust at high advance ratios, because the advancing side needs to be trimmed against the retreating side with significant reverse flow, in which the rotor is ineffective in generating thrust. Therefore, a hingeless rotor that allows the advancing side to generate more thrust can be rewarding in overall thrust potential. Wind tunnel tests were conducted up to an advance ratio of 0.7 to investigate the behavior of hingeless rotors at high advance ratios with lift offsets. Performance, control angles, hub vibratory loads and blade structural loads were compared with comprehensive analysis predictions from UMARC, plus the wing performance predictions from AVL. The results demonstrate that a hingeless rotor with lift offset is more efficient in generating thrust and exhibits higher lift-to-drag ratio at high advance ratios. The blade structural load level is significantly higher compared to an articulated rotor, especially for 2/rev flap bending moment, which can pose a critical structural constraint on the rotor.Item Effect of Interactional Aerodynamics on Computational Aeroacoustics of Sikorsky's Notional X2 Platform(2020) Bahr, Ian; Baeder, James; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)An in-house acoustics code, ACUM, was used in conjunction with full vehicle CFD/CSD coupling to create a computational aeroacoustic framework to investigate the effect of aerodynamic interactions on the acoustic prediction of a compound coaxial helicopter. The full vehicle CFD/CSD was accomplished by using a high- fidelity computational fluid dynamics framework, HPCMP CREATETM-AV Helios, combined with an in-house computational structural dynamics solver to simulate the helicopter in steady forward flight. A notional X2TD helicopter consisting of a coaxial rotor, airframe, and pusher propeller was used and split into three simulation cases: isolated coaxial and propeller, airframe and full helicopter configuration to investigate each component’s effect on the others noise as well as the total noise. The primary impact on the acoustic prediction was the inclusion of the airframe in the CFD simulation as it affected both coaxial rotors as well as the propeller. It was found that the propeller and coaxial rotors had a negligible impact on each other.Item CFD/CSD STUDY OF INTERACTIONAL AERODYNAMICS OF A COAXIAL COMPOUND HELICOPTER IN HIGH-SPEED FORWARD FLIGHT(2020) Klimchenko, Vera; Baeder, James; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)This work presents a computational study of the aerodynamic interactions that arise between the components of a high-speed lift-offset coaxial compound helicopter in forward flight. The objective of this study is to develop a computational methodology that would enable fundamental understanding of the complex aeromechanics of a modern lift-offset coaxial compound rotorcraft configuration in it's entirety. The modeling of a helicopter is a coupled aeroelastic problem, in which the aerodynamics is highly dependent on the structural dynamics, and vice versa. Therefore, the prediction of the rotorcraft airloads and blade deformations must be performed with sufficient fidelity to accurately model both aspects of the problem. A high-fidelity computational fluid dynamics framework, HPCMP CREATE$^{TM}$-AV Helios, was used in conjunction with an in-house comprehensive analysis solver, to simulate a lift-offset coaxial compound helicopter in forward flight. A notional X2TD helicopter consisting of a lift-offset coaxial rotor, airframe and an aft-mounted propeller, was modeled in this work. An in-house comprehensive analysis solver, PRASADUM, performed trim calculations and the structural modeling using low order aerodynamics. Conventionally, the comprehensive analysis rotor airloads that are computed from the built-in low order aerodynamic models, would be corrected with the high-fidelity CFD airloads using delta coupling procedure. In this study, the conventional rotor delta coupling methodology was used to study the interactional aerodynamics of a coaxial rotor system in forward flight at a range of flight speeds (50 knots to 225 knots). This study also focused on extending this methodology to perform high-fidelity airloads corrections for airframe and the propeller. The low order rotor, airframe and propeller aerodynamic loads were corrected with the high-fidelity CFD airloads, using a full vehicle loose delta coupling methodology. The two CFD/CSD coupling approaches, rotor and full vehicle, were compared. The results showed that correcting the low fidelity CSD airframe airloads with high-fidelity CFD airloads affects the rotor trim solution. The converged trim state from the full vehicle delta coupling procedure was utilized to study the fundamental interactional aerodynamics between various components of the coaxial compound helicopter. The CFD simulations were performed for isolated helicopter components and component combinations.Item An Experimental Investigation of Hypersonic Boundary-Layer Transition on Sharp and Blunt Slender Cones(2019) Kennedy, Richard Edward; Laurence, Stuart J; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)Understanding the instabilities leading to the laminar-to-turbulent transition of a hypersonic boundary layer is a key challenge remaining for the design of efficient hypersonic vehicles. In the present study, experiments are performed in three different facilities at freestream Mach numbers between 6 and 14 to characterize instability mechanisms leading to transition on a 7-degree half-angle slender cone. Second-mode instability waves are visualized using a high-speed schlieren setup with the camera frame rate and spatial resolution optimized to allow individual disturbances to be tracked. In order to facilitate quantitative time-resolved measurements, a method of calibrating the schlieren system and novel image-processing algorithms have been developed. Good agreement is observed between the schlieren measurements, surface pressure measurements, and parabolized stability equation computations of the second-mode most-amplified frequencies and N factors. The high-frequency-resolution schlieren signals enable a bispectral analysis that reveals phase locking of higher harmonic content leading to nonlinear wave development. Individual disturbances are characterized using the schlieren wall-normal information not available from surface measurements. Experiments are also performed to investigate the effect of nose-tip bluntness. For moderate to large bluntness nose tips, second-mode instability waves are no longer visible, and elongated structures associated with nonmodal growth appear in the visualizations. The nonmodal features exhibit strong content between the boundary-layer and entropy-layer edges and are steeply inclined downstream. Simultaneously acquired surface pressure measurements reveal high-frequency pressure oscillations typical of second-mode instability waves associated with the trailing edge of the nonmodal features.Item Gas Turbine / Solid Oxide Fuel Cell Hybrids: Investigation of Aerodynamic Challenges and Progress Towards a Bench-Scale Demonstrator(2019) Pratt, Lucas Merritt; Cadou, Christopher P; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)Modern aircraft are becoming more electric making the efficiency of on-board electric power generation more important than ever before. Previous work has shown that integrated gas turbine and solid oxide fuel cell systems (GT-SOFCs) can be more efficient alternatives to shaft-driven mechanical generators. This work advances the GT-SOFC concept in three areas: 1) It develops an improved model of additional aerodynamic losses in nacelle-based installations and shows that external aerodynamic drag is an important factor that must be accounted for in those scenarios. Additionally, this work furthers the development of a lab-scale prototype GT-SOFC demonstrator system by 2) characterizing the performance of a commercial off-the-shelf (COTS) SOFC auxiliary power unit that will become part of the prototype; and 3) combining a scaled-down SOFC subsystem model with an existing thermodynamic model of a small COTS gas turbine to create an initial design for the prototype.Item Computer Graphics Based Optical Tracking for Hypersonic Free-Flight Experiments(2019) Starshak, William; Laurence, Stuart; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)Aerodynamic measurement in hypersonic short-duration facilities – facilities with test times shorter than 10 milliseconds – is a topic of ongoing research. Standard force-balance approaches cannot handle the short test-time or the flow-initiating shock wave. Experimentalists have developed alternative techniques; but these techniques deliver merely adequate results at the cost of significant operational and – especially – calibration complexity. Recently, Laurence, et al. proposed using high-speed shadowgraph imaging and edge fitting (matching the visualized edge to an analytic equation for that edge) to make high-precision free-flight measurements of capsules. This new technique promised equivalent accuracy to existing techniques with far less pre-test calibration. The technique as developed, however, was limited to simple shapes in 2D motion. This thesis presents a generalization of the edge-fitting concept. Using the correspondence between a model's orientation and its silhouette, the trajectory of a model may be tracked to $1 \, \rm \mu m$ positional and $0.01^{\circ}$ angular accuracy. The silhouette is generated using computer-graphics techniques based upon a 3D mesh of the model's surface geometry. Consequently, the proposed technique is general to the model shape, the number of models, the properties of the camera imaging the experiment, and the number of cameras. Using the technique, we measured the hypersonic aerodynamics of a sphere, a blunt sphere-cone capsule, a lifting-body spacecraft, and the University of Maryland Testudo. In addition, multiple-camera and multiple-body tracking capability is demonstrated with an experiment investigating the dynamics of a breaking-up satellite. Results show that the method achieves accuracy comparable to or better than existing techniques with a simpler experimental procedure.
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