Theses and Dissertations from UMD
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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 Shock Train/Boundary-Layer Interaction in Rectangular Scramjet Isolators(2015) Geerts, Jonathan; Yu, Kenneth H; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)Numerous studies of the dual-mode scramjet isolator, a critical component in preventing inlet unstart and/or vehicle loss by containing a collection of flow disturbances called a shock train, have been performed since the dual-mode propulsion cycle was introduced in the 1960s. Low momentum corner flow and other three-dimensional effects inherent to rectangular isolators have, however, been largely ignored in experimental studies of the boundary layer separation driven isolator shock train dynamics. Furthermore, the use of two dimensional diagnostic techniques in past works, be it single-perspective line-of-sight schlieren/shadowgraphy or single axis wall pressure measurements, have been unable to resolve the three-dimensional flow features inside the rectangular isolator. These flow characteristics need to be thoroughly understood if robust dual-mode scramjet designs are to be fielded. The work presented in this thesis is focused on experimentally analyzing shock train/boundary layer interactions from multiple perspectives in aspect ratio 1.0, 3.0, and 6.0 rectangular isolators with inflow Mach numbers ranging from 2.4 to 2.7. Secondary steady-state Computational Fluid Dynamics studies are performed to compare to the experimental results and to provide additional perspectives of the flow field. Specific issues that remain unresolved after decades of isolator shock train studies that are addressed in this work include the three-dimensional formation of the isolator shock train front, the spatial and temporal low momentum corner flow separation scales, the transient behavior of shock train/boundary layer interaction at specific coordinates along the isolator's lateral axis, and effects of the rectangular geometry on semi-empirical relations for shock train length prediction. A novel multiplane shadowgraph technique is developed to resolve the structure of the shock train along both the minor and major duct axis simultaneously. It is shown that the shock train front is of a hybrid oblique/normal nature. Initial low momentum corner flow separation spawns the formation of oblique shock planes which interact and proceed toward the center flow region, becoming more normal in the process. The hybrid structure becomes more two-dimensional as aspect ratio is increased but corner flow separation precedes center flow separation on the order of 1 duct height for all aspect ratios considered. Additional instantaneous oil flow surface visualization shows the symmetry of the three-dimensional shock train front around the lower wall centerline. Quantitative synthetic schlieren visualization shows the density gradient magnitude approximately double between the corner oblique and center flow normal structures. Fast response pressure measurements acquired near the corner region of the duct show preliminary separation in the outer regions preceding centerline separation on the order of 2 seconds. Non-intrusive Focusing Schlieren Deflectometry Velocimeter measurements reveal that both shock train oscillation frequency and velocity component decrease as measurements are taken away from centerline and towards the side-wall region, along with confirming the more two dimensional shock train front approximation for higher aspect ratios. An updated modification to Waltrup \& Billig's original semi-empirical shock train length relation for circular ducts based on centerline pressure measurements is introduced to account for rectangular isolator aspect ratio, upstream corner separation length scale, and major- and minor-axis boundary layer momentum thickness asymmetry. The latter is derived both experimentally and computationally and it is shown that the major-axis (side-wall) boundary layer has lower momentum thickness compared to the minor-axis (nozzle bounded) boundary layer, making it more separable. Furthermore, it is shown that the updated correlation drastically improves shock train length prediction capabilities in higher aspect ratio isolators. This thesis suggests that performance analysis of rectangular confined supersonic flow fields can no longer be based on observations and measurements obtained along a single axis alone. Knowledge gained by the work performed in this study will allow for the development of more robust shock train leading edge detection techniques and isolator designs which can greatly mitigate the risk of inlet unstart and/or vehicle loss in flight.Item Effect of Fin-Guided Fuel Injection on Supersonic Mixing and Combustion(2014) Aguilera Munoz, Camilo; Yu, Kenneth H; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)Rapid mixing and combustion is a key challenge in supersonic combustors due to the extremely short flow residence time and the effect of compressibility. Mixing enhancement is therefore desirable to ensure timely mixing, reaction, and heat release. Fin-guided fuel injection is one approach that can be optimized for propulsion performance consideration. The present investigation examined the mixing and combustion characteristics of using this alternative fuel injection method to evaluate its performance in comparison to conventional transverse wall injection. This study was conducted in two parts: (1) fuel-air mixing experiments in a non-reacting Mach 2.2 flow with a test section Reynolds number of 1.15×106, and (2) combustion experiments using a high-enthalpy, vitiated air flow with a Mach 2.0 condition at the isolator inlet and Reynolds number of 1.14× 105. The non-reacting mixing study used either helium or ethanol, while the combustion study used either hydrogen or ethylene as fuel for each experiment. The mixing behavior of the gaseous and liquid jets was studied using schlieren and a laser sheet technique while quantitative assessments were made from pressure measurements. Similarly, the physical mechanisms in the reacting flow experiments were analyzed using schlieren visualizations while pressure measurements and chemiluminescence emission data were used for performance evaluation. The fuel-air mixing study highlighted possible tradeoffs between mixing enhancement and the stagnation pressure loss stemming from fuel jet-induced shocks. Since the fin was designed to weaken the oblique shock strength while shielding the fuel jet penetrating into the core airflow, it not only resulted in better mixing but also improved the pressure recovery. For gaseous fuel, fin-guided injection improved jet penetration by 100 to 200% for a momentum ratio between 0.15 and 0.03. It also resulted in 64 to 85% additional pressure recovery of the injection shock loss. Combustion experiments revealed that the fin could be used to extend the upper limit of supersonic combustion mode in the present configuration, from an equivalence ratio of 0.04 to 0.12, by preventing thermal choking caused by concentrated heat release near the baseline flame holder. This could be advantageous for certain systems by reducing the thermal protection requirements. However, the fin also made the wall cavity flame holder less effective by increasing fuel penetration away from the bottom wall. The net effects on propulsion system performance will ultimately depend on whether ramjet or scramjet mode is preferred for a given operation.Item ENTROPY CONSIDERATIONS APPLIED TO SHOCK UNSTEADINESS IN HYPERSONIC INLETS.(2012) Bussey, Gillian Mary Harding; Lewis, Mark J; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)The stability of curved or rectangular shocks in hypersonic inlets in reponse to flow perturbations can be determined analytically from the principle of minimum entropy. Unsteady shock wave motion can have a significant effect on the flow in a hypersonic inlet or combustor. According to the principle of minimum entropy, a stable thermodynamic state is one with the lowest entropy gain. A model based on piston theory and its limits has been developed for applying the principle of minimum entropy to quasi-steady flow. Relations are derived for analyzing the time-averaged entropy gain flux across a shock for quasi-steady perturbations in atmospheric conditions and angle as a perturbation in entropy gain flux from the steady state. Initial results from sweeping a wedge at Mach 10 through several degrees in AEDC's Tunnel 9 indicates the bow shock becomes unsteady near the predicted normal Mach number. Several curved shocks of varying curvature are compared to a straight shock with the same mean normal Mach number, pressure ratio, or temperature ratio. The present work provides analysis and guidelines for designing an inlet robust to off- design flight or perturbations in flow conditions an inlet is likely to face. It also suggests that inlets with curved shocks are less robust to off-design flight than those with straight shocks such as rectangular inlets. Relations for evaluating entropy perturbations for highly unsteady flow across a shock and limits on their use were also developed. The normal Mach number at which a shock could be stable to high frequency upstream perturbations increases as the speed of the shock motion increases and slightly decreases as the perturbation size increases. The present work advances the principle of minimum entropy theory by providing additional validity for using the theory for time-varying flows and applying it to shocks, specifically those in inlets. While this analytic tool is applied in the present work for evaluating the stability of shocks in hypersonic inlets, it can be used for an arbitrary application with a shock.