UMD Theses and Dissertations
Permanent URI for this collectionhttp://hdl.handle.net/1903/3
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 given thesis/dissertation in DRUM.
More information is available at Theses and Dissertations at University of Maryland Libraries.
Browse
11 results
Search Results
Item EXPERIMENTAL INVESTIGATION OF BOUNDARY LAYER TRANSITION ON CONE-FLARE GEOMETRIES AT MACH 4(2024) Norris, Gavin; Laurence, Stuart J; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)This study investigates supersonic boundary layer transition on a cone-flarewith a 5° half-angle straight cone and flared bases of +5°, +10°, and +15°. The experiments used the University of Maryland's Multiphase Flow Investigations Tunnel (MIST), a Mach 4 Ludweig tube. Experiments were performed “dry”, without aerosols or droplets, and focus on the first-mode (Tollmien-Schlichting) boundary layer instability waves and their interaction with the compression corner. Using high-speed Schlieren imaging, the boundary layer dynamics on the cone-flare's top surface were analyzed. The data were processed through Power Spectral Density (PSD) and Spectral Proper Orthogonal Decomposition (SPOD) techniques to study the behavior of the first-mode waves and the transition location changes. The findings reveal coherent wave packets within the boundary layer at frequencies characteristic of the first-mode. The wave packets power increased along the cone and peaked near the compression corner before dissipation on the flare. These findings contribute to the understanding of first-mode boundary layer transition mechanisms in hypersonic flows for the cone-flare geometry.Item Aero Database Development and Two-Dimensional Hypersonic Trajectory Optization for the High-speed Army Reference Vehicle(2023) James, Brendan; Brehm, Christoph; Larsson, Johan; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)Steady-flow inviscid and simulations of the High-Speed Army Reference Vehicle geometry were performed within the CHAMPS solver framework at Mach numbers of 4, 6, and 8, and an integrated streamline method was used to apply viscous corrections for Reynolds numbers up to 2x10^8. For each flow Mach, angle of attack sweeps from -10° to +10° were used to determine baseline drag, lift, and moment coefficient alpha dependencies. Coefficient values were then interpolated across Mach, alpha, and Reynolds number parameter spaces to construct an aerodynamic force coefficient database for use in two-dimensional flight simulation and trajectory optimization. By simulating flight with a maximum lift-to-drag control input, sample trajectories for determining maximum vehicle range were produced. A proportional-navigation (PN) controller was implemented which allowed for the targeting of specific altitudes throughout the progression of a trajectory. The PN controller and simulation schemes were then utilized in genetic-algorithm optimization to produce trajectory profiles for achieving minimum time-to-target for gliding flight in standard atmospheric conditions. Over the examined range of initial altitudes, Mach numbers, and release angles, the fastest trajectories were consistently shown to be those which achieved or maintained stratospheric altitudes and consequently benefited from significantly reduced drag before performing a nose-over maneuver for an accurate ground strike.Item SCRAMJET COMBUSTOR MODE TRANSITION BY CONTROLLING FUEL INJECTION DISTRIBUTION(2022) Kanapathipillai, Mithuun; Yu, Kenneth H; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)Dual-mode scramjets are able to expand the operable Mach number range of the simplescramjet through manipulation of a thermal throat. Using the thermal throat, the scramjet can operate in either thermally-choked mode or supersonic combustion mode. The transition between these two events is still not very well understood. Past research has shown natural combustor mode transition to be highly unstable and characterized by frequent mode hopping. Long timescales associated with combustor mode transition also increase the potential for combustion dynamic events to occur. Due to the high level of hysteresis present in these events, designing a way to precisely controlling mode transition timing proves to be an ongoing challenge. The present research seeks to use a distributed fuel injection method to control and better understand combustor mode transition behavior. This study was performed using a laboratory-scale, direct-connect scramjet combustor. The facility simulated Mach 5 flight conditions using vitiation to match the enthalpy conditions necessary and to recreate typical isolator and combustor flowfields characteristic of the dual-mode scramjet. A supersonic nozzle was employed to achieve an isolator inlet Mach number of 2.0. For the reacting flow tests, gaseous hydrogen was injected through one to four injectors using a distributed fuel injection scheme while keeping a global equivalence ratio of 0.52 constant. Various imaging diagnostics and wall pressure measurements were used to better study the relationship between combustor behavior and the number of fuel injectors. The findings revealed that combustor mode operation has a significant effect on combustor performance, as indicated by the pressure rise, axial heat release distributions, and local flowpath Mach number for these cases. The deduced heat release for the single injection case showed that most of the heat release occurs near the cavity flame-holder leading to a relatively large pressure jump causing a premature transition to thermal choking. In the case of distributed injection, heat release occurs more evenly across the expanding portion of the combustor, which prevents early transition to thermal choking. Active control of the combustor mode transition event is demonstrated through the use of a fuel injection distribution and scheduling system as well as fast-response solenoid valves. In the active control cases, the global equivalence ratio was maintained at 0.52, but using the active control system the combustor was able to bidirectionally switch between stable, thermally-choked mode and stable, scramjet mode. Furthermore, actively controlled mode transition occurs at much faster timescales than what was observed for natural mode transition. This allows for the potential to actively control on demand combustor mode transition in a real world dual-mode ramjet-scramjet combustor through appropriately scheduled fuel injection distribution.Item DEVELOPMENT OF KRYPTON PLANAR LASER INDUCED FLUORESCENCE METHODS FOR THE MEASUREMENT OF HYPERSONIC FLOW CONDITIONS(2021) Standage, Chase; Gupta, Ashwani; Yu, Kenneth; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)Conventional wind tunnel flow measurement techniques typically involve the use of intrusive sensor systems, such as Pitot-probes and transducers which come in contact with the flow. Intrusive methods become impractical for high Mach number flows, as such methods can cause considerable disruption to the integrity of flow measurements. Therefore, it is desirable to utilize non-intrusive methods in such experiments, especially as hypersonic flow conditions are achieved. Schlieren and shadowgraph imaging methods have been used successfully for decades as a method of non-intrusive flow visualization. However, these methods become obsolete when the path of light is obstructed, which is a common problem when analyzing concave surfaces and complex geometries. The goal of this project was to develop a scalable krypton planar laser induced fluorescence flow visualization system for use on curved-surface geometries in sup- port of the hypersonic Boundary Layer Transition (BOLT) program. The system was designed to fit multiple wind-tunnel facilities, including the AEDC Tunnel 9 hypersonic test facility and UMD Ludwieg Tube. In order to design and test the system, the AEDC Mach 3 Calibration wind tunnel was utilized and Kr-PLIF measurements were taken about a 0.50” spherical model and 2” BOLT model. A wide variety of equipment and methods were assessed for their suitability of this project, including 3 cameras and 7 sheet combinations. A beam from a single-diode Ti-Sapphire laser was amplified, modulated, and shaped in order to create a thin laser-sheet of 0.25-1.0” width and 0.01”-0.025” thickness, frequency of 1 kHz, and pulse width of 40 fs. The flow was seeded with 5% krypton, and tests were conducted at Mach 3. The results were compared to Schlieren imaging tests conducted onsite in the same Mach 3 wind tunnel. The Kr-PLIF method was moderately successful in finding regions of relatively high flow density, such as boundary layers and leading edges at an angle-of-attack. Additionally, Kr-PLIF was able to make measurements about the curved region of the BOLT model, which was previously unobservable by Schlieren imaging.Item Response of hypersonic boundary-layer disturbances to compression and expansion corners(2021) Butler, Cameron Scott; Laurence, Stuart; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)An experimental campaign was conducted at the University of Maryland - College Park to examine the impact of abrupt changes in surface geometry on hypersonic boundary-layer instability waves. A model consisting of a 5-degree conical forebody was selected to encourage the dominance of second-mode wavepackets upstream of the interaction region. Interchangeable afterbody attachments corresponding to flow deflections of -5-degree to +15-degree in 5-degree increments were considered. The adverse pressure gradient imposed by the +10-degree and +15-degree configurations caused the boundary layer to separate upstream, creating a region of recirculating flow. High-speed schlieren (440-822 kHz) was employed as the primary means of flow interrogation, with supplemental surface measurements provided by PCB132B38 pressure transducers. A lens calibration was applied to the images to provide quantitative fluctuations in density gradient. The high frame rate made possible the use of spectral analysis techniques throughout the entire field of view. This analysis reveals complex growth and decay trends for incoming second-mode disturbances. Additional, low-frequency content is generated by the deflected configurations. This is most pronounced for the separated cases where distinct, shear-generated disturbances are observed. Spectral proper orthogonal decomposition (SPOD) is demonstrated as a powerful tool for resolving the flow structures tied to amplifying frequencies. Nonlinear interactions are probed through bispectral analysis. Resonance of low-frequency structures is found to play a large role in nonlinear energy transfer downstream of the compression corners, particularly for the separated cases. Concave streamline curvature appears to result in concentrated regions of increased nonlinearity. These nonlinear interactions are shown to be spatially correlated with coherent flow structures resolved through SPOD. Finally, a limited computational study is carried out to demonstrate the ability of linear stability theory and the parabolized stability equations to reproduce experimental results obtained for the +10-degree extension. The development of the second-mode and shear-generated disturbances resolved by the computational analysis shows excellent agreement with the experimental results.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.Item Morphing Waveriders for Atmospheric Entry(2019) Maxwell, Jesse R; Oran, Elaine S; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)The primary challenge for vehicles entering planetary atmospheres is surviving the intense heating and deceleration encountered during the entry process. Entry capsules use sacrificial ablative heat shields and sustain several g deceleration. The high lift produced by the Space Shuttle geometry resulted in lower rates of heating and deceleration. This enabled a fully reusable vehicle that was protected by heat shield tiles. Hypersonic waveriders are vehicles that conform to the shape of the shock wave created by the vehicle. This produces high compression-lift and low drag, but only around a design Mach number. Atmospheric entry can reach speeds from zero to as high as Mach 40. A morphing waverider is a vehicle that deflects its flexible bottom surface as a function of Mach number in order to preserve a desired shock wave shape. It was demonstrated in this work that doing so retains high aerodynamic lift and lift-to-drag ratio across a wide range of Mach number. Numerical simulations were conducted for case-study waveriders designed for Mach 6 and 8 for flight at their design conditions as well as with variations in angle-of-attack and Mach number. A single-species air model was used between Mach 1 and 12 with the RANS k-omega SST and LES-WALE turbulence models. A seven-species air model was used for Mach 15 at 60km altitude and Mach 20 at 75km. Analytical methods were used to construct a reduced-order model (ROM) for estimating waverider aerodynamic forces, moments, and heating. The ROM matched numerical simulation results within 5-10% for morphing waveriders with variations in angle-of-attack, but discrepancies exceeded 20% for large deviations of rigid vehicles from their design Mach numbers. Atmospheric entry trajectory simulations were conducted using reduced-order models for morphing waverider aerodynamics, the Mars Science Laboratory (MSL) capsule, and the Space Shuttle. Three morphing waveriders were compared to the Space Shuttle, which resulted in reduced heating and peak deceleration. One morphing waverider was compared to the MSL capsule, which demonstrated a reduction in the peak stagnation heat flux, a reduction in the peak and average deceleration, and a reduction in the peak area-averaged heating.Item Development and Application of Mach 10 PIV in a Large Scale Wind Tunnel(2018) Brooks, Jonathan; Gupta, Ashwani K; Marineau, Eric C; Mechanical Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)This dissertation presents the development of particle image velocimetry (PIV) for use in a large-scale hypersonic wind tunnel to measure the turbulent boundary layer (TBL) and shock turbulent boundary layer interaction (STBLI) on a large hollow cylinder flare (HCF) test article. The main feature of this application of PIV is the novel local injector which injects seeding particles into the high-speed section of the flow. Development work began sub-scale in a Mach 3 wind tunnel where the seeding particle response was characterized and the local injectors were demonstrated. Once the measurement technique was refined, it was scaled up to hypersonic flow. The particle response was characterized through PIV measurements of Mach 3 TBLs under low Reynolds number conditions, $ Re_\tau=200{-}1,000 $. Effects of Reynolds number, particle response and boundary layer thickness were evaluated separately from facility specific experimental apparatus or methods. Prior to the current study, no detailed experimental study characterizing the effect of Stokes number on attenuating wall normal fluctuating velocities has been performed. Also, particle lag and spatial resolution are shown to act as low pass filters on the fluctuating velocity power spectral densities which limit the measurable energy content. High-speed local seeding particle injection has been demonstrated successfully for the first time. Prior to these measurements, PIV applications have employed global seeding or local seeding in the subsonic portion of the nozzle. The high-speed local seeding injectors accelerate the particle aerosol through a converging/diverging supersonic nozzle which exits tangentially to the wall. Two methods are used to measure the particle concentration which shows good agreement to the CFD particle tracking codes used to design the injector nozzle profiles. Based on the particle concentration distribution in the boundary layer a new phenomenon of particle biasing has been identified and characterized. PIV measurements of a Mach 10 TBL and STBLI have been performed on a large (2.4-m long, 0.23-m dia.) HCF at a freestream unit Reynolds number of 16 million per meter. These are the highest Mach number PIV measurements reported in the literature. Particles are locally injected from the leading edge of the test article and turbulent mixing dispersed the particles for a relatively uniform high concentration of particles at the measurement section 1.83-m downstream of the leading edge. The van Driest transformed mean velocity in the TBL agrees well with incompressible zero pressure gradient log law theory. Morkovin-scaled streamwise velocity fluctuations agree well with the literature for the majority of the boundary layer.Item ROBUST MULTI-OBJECTIVE OPTIMIZATION OF HYPERSONIC VEHICLES UNDER ASYMMETRIC ROUGHNESS-INDUCED BOUNDARY-LAYER TRANSITION(2014) Ryan, Kevin Michael; Lewis, Mark J; Yu, Kenneth H; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)The effects of aerodynamic asymmetries on hypersonic vehicle controllability and performance were investigated for a wide range of geometries. Asymmetric conditions were introduced by an isolated surface roughness that forces boundary-layer transition resulting in a turbulence wedge downstream of the disturbance. The disturbance simulates the effects of physical deformations that may exist on a vehicle surface or leading edge, such as protruding edges of thermal protection system tiles or non-uniform surface roughness. Both multi-objective and robust multi-objective optimization studies were performed. Traditional multi-objective optimization methods were used to identify vehicle designs that are best suited to withstand spanwise asymmetric boundary-layer transition while retaining its performance and payload requirements. Trade-offs between vehicle controllability and performance were analyzed. A novel multi-objective based robust optimization method to solve single-objective optimization problems with environmental parameter uncertainty was proposed and tested. Unlike commonly used robust optimization methods, the multi-objective method formulates an optimization problem such that post-optimality data handling techniques can identify multiple robust designs from a single solution set. This allows for comparisons to be made between different types of robust designs, thus providing more information about the design space. Comparisons were made between the robust multi-objective optimization formulation and conventional robust regularization- and aggregation-based methods. The results, performance, and philosophies of each method are discussed. Design trends were identified for classifying the optimum and robust optimum designs of hypersonic vehicle shapes under boundary-layer transition uncertainties. Traditional multi-objective optimization results show that two types of vehicle shapes bound the set of Pareto-optimal solutions: wedge-like and cone-like. The L2-norm optimum design, representing a compromise between the competing shapes, was a hybrid wedge-cone shape. The robust optimization results show that a flat wedge-like vehicle design is best for a worst-case scenario, while a pyramidal shaped vehicle design minimizes the expected detrimental effects on vehicle controllability. The analyses prove that the novel robust optimization method can provide a range of robust optimum results, while also capturing trade-offs within the design space, providing capabilities not available in state-of-the-art robust optimization methods.Item Hypersonic Application of Focused Schlieren and Deflectometry(2010) VanDercreek, Colin Paul; Yu, Kenneth H; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)A non-intrusive diagnostic capability for determining the hypersonic shock and boundary layer structure was developed, installed, and successfully tested at the AEDC Hypervelocity Tunnel 9. This customized diagnostic involves a combination of a focused schlieren system, which relies on creating multiple virtual light sources using a Fresnel lens and a source grid, and a deflectometry system, which uses the focused schlieren and a photomultiplier tube. It was used for obtaining spatially resolved images of density gradients with a depth of focus less than one centimeter, while allowing high frequency measurements of density fluctuations. The diagnostic was applied in investigating the second-mode instability waves present in the boundary layer of a sharp-nosed cone submerged in a Mach 10 flow. The waves were successfully imaged and their frequencies were measured even though the flow density was below 0.01 kg/m^3 and the frequencies over 200 kHz. This adds a new capability to hypersonic testing.