A direct connect, supersonic solid fuel combustor with a cavity is explored in the context of understanding characteristics related to ignition, regression rate, combustion, and flow fields for application in advancing solid fuel scramjet research. 3D printed, polymethylmethacrylate fuel grains are loaded into both fully enclosed and optically accessible combustors. The ignition characteristics are investigated by systematically varying the internal geometry of the fuel grain to develop a flammability map with respect to non-dimensional geometric parameters. Results reveal that a longer and larger flameholding cavity creates favorable conditions for ignition and sustained combustion. The inlet temperature is also systematically varied to extend the available literature on the supersonic combustion of solid fuels to lower temperature operating conditions and show that a higher inlet temperature is conducive to sustained combustion and higher regression rates. The regression rates of the fuel grains are measured to determine a concentration of regression in the flameholding cavity along the angle of the downstream side of the cavity. Ignition and sustained combustion rely heavily on the fuel in the flameholding cavity. A decreasing regression rate is observed as the fuel regresses by measuring the regression rate at discrete time intervals during a firing of the optical combustor. The optical combustor is also subject to various high-frequency imaging techniques. Shadowgraph imaging shows the changes in density of the flow field and finds a normal shock in the constant area section. CH* chemiluminescence imaging provides novel observations of the concentrated areas of combustion along the fuel grain wall by highlighting the heat release from combustion. A high intensity of CH* radicals is in the upstream section of the flameholding cavity. When considered in the context of the concentration of regression, this indicates that the recirculation zone pulls fuel from the downstream section of the cavity, combusts it in the upstream section of the flameholding cavity, then expels the higher enthalpy gas into the core flow. Additionally, observing the flow provides insight into the flow dynamics of opposing cavities in a supersonic flow field. The symmetry of the flow field is found to be reliant on the stability of the flameholding cavity length to depth ratio. / Master of Science / A solid fuel scramjet has the potential to be the simplest and most cost effective method of achieving hypersonic flight. A liquid fuel scramjet has been demonstrated in free flight, but liquid fuels present many issues involving safety and storage that can be eliminated by introducing solid fuels. Supersonic combustion, or burning fuel in an air flow moving faster than the speed of sound, is a complicated subject due to the irregularity of flow fields and the requirement of combustion to occur at a high rate. The research within this thesis presents many novel technologies that have never been presented in published literature in the context of the supersonic combustion of solid fuels. By conducting ground testing of a solid fuel scramjet, characteristics of the combustion can be studied to expand the available literature in the field to new fuel geometries and inlet conditions. The ignition and sustained combustion of a solid fuel scramjet is extremely reliant on the initial geometry of the fuel and the initial temperature of the flow. This research advances the field of supersonic combustion of solid fuels by developing an optically accessible combustor using quartz windows. These characteristics of supersonic combustion are investigated using highspeed video recording. The results of these techniques provide insight into favorable fuel geometries and inlet conditions. Additionally, patterns observed in the flow field explain concentrations of combustion and fuel consumption.

<p>Air intakes are a fundamental part of all high speed airbreathing propulsion concepts. The main purpose of an intake is to capture and compress freestream air for the engine. At hypersonic speeds, the intake’s surface and shock structure effectively slow the airflow through ram-air compression. In supersonic-combustion ramjets, the captured airflow remains supersonic and generates complicated shock structures. The design of these systems require careful evaluation of proposed operating conditions and relevant aerodynamic phenomena. The physics of these systems, such as the intake’s operability range, mass capture efficiency, back-pressure resiliency, and intake unstart margins are all open areas of research. </p> <p><br></p> <p>A high speed intake, dubbed the Indiana Intake Testbed, was developed for experimentation within the Boeing-AFOSR Mach 6 Quiet Tunnel at Purdue University. This inward-turning, mixed compression intake was developed from osculating axisymmetric theory and uses a streamtracing routine to create a shape-transitioned geometry. To account for boundary layer growth, a viscous correction was implemented on the intake’s compression surfaces. This comprehensive independent design code was pursued to generate an unrestricted geometry that satisfies academic inquiry into fluid dynamic interactions relevant to intakes. Additionally, the design code contains built-in analysis tools that are compared against CFD calculations and experimental data. </p> <p><br></p> <p>Two blockage models were constructed and outfitted with Kulite pressure transducers to detect possible intake start and unstart effects. Due to an error in the design code, the preliminary blockage models’ lower surfaces were oversized. The two intake models were tested over a freestream Reynolds number sweep, under noisy and quiet flow, at one non-zero angle of attack, and at a singular back-pressure condition. Back-pressure effects acted to unstart the intake and provide a comparison between forced-unstart and started states. The experimental campaign cataloged both tunnel starting and inlet starting conditions, which informed the design of the finalized model. The finalized model is presented herein. Future experiments to study isolator shock-trains, shock-wave boundary layer interactions, and possible instances of boundary layer transition on the intake’s compression surface are planned. </p>

<p>In rotating detonation engines the turbine inlet conditions may be transonic with unprecedented unsteady fluctuations, very different from those in conventional high-pressure turbines. To ensure an acceptable engine performance, the turbine passages must be unchoked at subsonic and started at supersonic conditions. Additionally, to maximize the aerodynamic performance potential, ad-hoc designs are required, suited for the oscillations in Mach number and flow angle. This manuscript focuses on designing and characterizing diffusive turbine vanes that can operate downstream of a transonic rotating detonation combustor.  </p> <p><br></p> <p>First, the phenomenon of unstarting is presented, concentrating on the effect of pressure loss on the accurate prediction of the starting limit. Afterward, a multi-objective optimization with steady Reynolds Averaged Navier Stokes simulations, including the endwall and 3D vane design, is performed. The results are discussed, highlighting the impact of the throat-to-inlet area ratio on the pressure loss and the geometric features of the top-performing designs. Compared to previous  research on stator passages with contoured endwalls, considerable reductions in pressure loss and stator-induced rotor forcing are obtained, with an extended operating range and preserving high turning.  </p> <p><br></p> <p>Subsequently, the influence of the inlet boundary layer thickness on the vane performance is evaluated, inducing remarkable increases in pressure loss and downstream pressure distortion. Employing an optimization with a thicker inlet boundary layer, specific endwall design recommendations are found, providing a notable improvement in both objective functions. The impact of the geometry variations on flow detachment is assessed as well.</p> <p><br></p> <p>Finally, the impact of the inlet flow angle on the vane design is studied through a multi-point, multi-objective optimization with different inlet angles. The effect of incidence on the flow field and vane performance is evaluated first. Then, by comparing the optimized geometries with those optimized for axial inflows, several design guidelines are identified </p>