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The oblique and traverse configurations of injecting gaseous fuel in a low-aspect-ratio supersonic combustor are characterized and compared numerically using a validated code. Non-reacting conditions are considered, where fuel is simulated by helium. The combustor, which has a rectangular cross-section of constant span, is attached to a Mach 2 nozzle and expands along the top and bottom walls. A choked wall port is used for both injection configurations. Different sets of operating conditions have been simulated. It was found that injecting fuel obliquely results in higher efficiency as well as effectiveness. Unlike the traverse configuration, oblique injection makes use of the beneficial interaction of the injection-induced shock waves with the air/fuel shear layer. This interaction was proven in previous research to be effective for mixing enhancement in supersonic flows. However, in contrast with the results of previous research, normal or oblique injection at large angles (30° or 60°) is not necessary for the achievement of sufficient mixing in supersonic flows. Substantial mixing improvement was found at angles as small as 5°. Fuel injection at such small angles improves the fuel-air mixing while minimizing the injection-induced pressure losses, which leads to increased thrust. Our results on mixing under non-reacting conditions provide good preliminary insights on a more favorable fuel injection configuration that provides better mixing with lower losses and higher thrust.

A thin wedge-shaped blade, sometimes known as pylon, was used to enhance transverse fuel mixing in Mach 2 airflow. Supersonic mixing experiments were conducted using 45° and 90° fuel injections from the wall in the immediate wake of the pylon blade, and the results were compared with a baseline case using 90° transverse fuel injection without any pylon assistance. All the injectors had a same-size diameter, and the injector diameter matched the maximum blade thickness at the pylon base. For both qualitative and quantitative comparison, wall-pressure measurements, planar Mie-scattering of smokeseeded fuel streams, and instantaneous and time-averaged schlieren visualization were employed. Fuel penetration height was measured as a function of axial distance from the schlieren images, while flow losses associated with pylon and jet-induced shocks were assessed from the wall pressure measurements. Also, the wake flow extending downstream from the pylon base was characterized by planar Mie-scattering images. Substantial improvement in mixing performance was observed with the use of pylon as the fuel penetration height was increased by 100~120% and the flow losses associated with jetinduced shock were reduced by 13~30%. The 90° injection achieved greater fuel penetration heights, while the 45° injection incurred greater savings in flow losses. The results open up the possibility of further increasing the performance by optimizing the fuel injection angle behind the pylon.

Nonreacting supersonic mixing experiments were conducted to characterize and compare two different fuel-injection schemes applied to a high-aspect ratio supersonic combustor. The combustor, which has a constant-width rectangular cross-section, is attached to a Mach 2 nozzle and expands along the top and bottom walls. Helium was injected through a choked orifice using transverse wall injection and parallel ramp injection. Pressure traces along the top wall of the combustor were used to determine flow field characteristics as well as to evaluate the static pressure rise caused by each system. Instantaneous and time-averaged Schlieren images were taken to visualize the resulting flow structures of each injection and to measure fuel penetration depth. Planar Mie scattering using ethanol droplets was also conducted to visualize the lateral and vertical spreading of the fuel in the near and far field. Results show that the ramp can be used with parallel injection providing just as effective fuel penetration as in the normal injection, but it also caused higher static pressure rise indicating increased flow loss in the main airflow. While combustion experiments are necessary to assess the impact on actual combustion performance and thermal efficiency, the present nonreacting mixing study provides more insights into the limitations of each injector configuration.

An experimental study of a reacting particle-laden flow has been conducted. An optically accessible two-dimensional combustor has been designed and fabricated to investigate the potential of aluminum nanoparticles as additives for fuels in airbreathing propulsion systems. Experiments were conducted with hydrocarbon only (ethylene) and mixtures of ethylene and aluminum. The oxidation of aluminum has been studied through the emission of AlO. Planar Mie Scattering was employed to ensure that uniform dispersion of the particles was achieved. Temperature measurements inside the combustor using thermocouples have been made in order to estimate the efficiency of combustion for a variety of equivalence ratios ranging from 0.52 to 0.7, aluminum loadings ranging from 8.7 to 18.1, and combustor residence times ranging from 6 to 10.5 milliseconds. Combustion efficiencies in aluminized and nonaluminized tests were found to be similar within experimental error and in all cases efficiencies greater than 90% were measured. The results open up the possibility of using nano-aluminum additives in fuels for airbreathing propulsion. Nomenclature

An experimental study to evaluate the potential of nano-sized boron as a fuel additive for high-speed airbreathing propulsion has been conducted. Experiments were conducted with hydrocarbon only (ethylene) and mixtures of ethylene and boron. The oxidation of boron was studied through the emission of BO 2. Temperature measurements inside the combustor using thermocouples have been made in order to determine boundaries in which the addition of boron provides a positive thermal output for a variety of equivalence ratios ranging from 0.54 to 0.69, boron loadings ranging from 9.7 to 15.2% by weight, and combustor residence times ranging from 6 to 10.5 milliseconds (combustor inlet velocities of 40 – 70 m/s). In all cases, an ethylene/oxygen pilot flame provided successful ignition of the boron particles. However, both BO 2 emission data and temperature measurements indicate that a critical temperature exists for sustained combustion of the particles. Tests with measured peak temperatures below 1700 K indicated no benefit of boron addition. Tests with measured peak temperatures above 1770 K showed a positive thermal contribution from boron addition. These results suggest that even employing nano-sized boron, only a small envelope for complete energetic extraction exists.

An experimental study of a three-dimensional Mach 2 scramjet combustor has been conducted. The combustor featured a square cross-section and a three-dimensional expanding section, with optical access on one side. A gaseous ethylene and gaseous oxygen pilot flame was used to ignite the primary fuel in the form of gaseous ethylene (C2H4) at various core air mass flow rates. The primary fuel was injected perpendicular from a combustor wall at just downstream of the pilot flame. Wall pressure measurements were taken along the length of the combustor. The measurements revealed that supersonic combustion was achieved for a limited axial distance downstream of the fuel injection. High- speed flame imaging was used to help characterize the supersonic reaction zone. In addition, C2* and CH* Chemiluminescence data were obtained for a range of equivalence ratios in order to determine heat release patterns. Relative combustion performance was assessed for each case by comparing the wall pressure traces and analyzing the flame visualization data. The results showed that the maximum performance was reached when the overall equivalence ratio was only 0.25 due to poor mixing. This study aims to provide the baseline performance data using normal injection, which will be used to compare combustor performance using different geometries and different injection schemes.