Research was conducted at Virginia Tech on a high-pressure uncooled plasma torch to study torch operational characteristics with hydrocarbon feedstocks and to determine the feasibility of using the torch as an igniter in scramjet applications. Operational characteristics studied included electrical properties, such as arc stability, voltage-current characteristics and start/re-start capabilities, and mechanical properties, such as coking, electrode erosion and transient to steady-state torch body temperature trends. Possible use of the plasma torch as an igniter in high-speed combustion environments was investigated through the use of emission spectroscopy and a NASA chemical kinetics code. All feedstocks tested; argon, methane, ethylene and propylene, were able to start. The voltage data indicated that there were two preferred operating modes, which were well defined for methane. For all gases, a higher current setting, on the order of 40 A, led to more stable torch operation. A low intensity, high frequency current applied to the torch, along with the primary DC current, resulted in virtual elimination of soot deposits on the anodes. Electrode erosion was found to multiply each time the complexity of the hydrocarbon was increased. Audio and high-speed visual analysis led to identification of 180 Hz plasma formation cycle, related to the three-phase power supply. The spectroscopic analysis aided in the identification of combustion enhancing radicals being produced by the torch, and results of the chemical kinetics analysis verified combustion enhancement and radical production through the use of a basic plasma model. Overall, the results of this study indicate that the plasma torch is a promising source for scramjet ignition, and further study is warranted. / Master of Science

Though the field of hypersonic vehicle design is thriving again, few studies to date demonstrate the technology through a mission in which multiple flight conditions and constraints are encountered. This is likely due to the highly integrated and sensitive nature of hypersonic vehicle components. Consequently, a formal Mach 6 through Mach 10 flight envelope is explored which includes cruise, acceleration/climb, deceleration/descend and turn mission segments. An exergy approach to the vehicle synthesis/design, in which trade-offs between dissimilar technologies are observed, is proposed and measured against traditional methods of assessing highly integrated systems. A quasi one-dimensional hypersonic vehicle system simulation program was constructed. Composed of two sub-systems, propulsion and airframe, mechanisms for loss are computed from such irreversible processes as shocks, friction, heat transfer, mixing, and incomplete combustion. The propulsion sub-system consists of inlet, combustor, and nozzle, while the airframe provides trim and force accounting measures. An energy addition mechanism, based on the potential of MHD technology, is utilized to maintain a shock-on-lip inlet operating condition. Thirteen decision variables (seven design and six operational) were chosen to govern the vehicle geometry and performance. A genetic algorithm was used to evaluate the optimal vehicle synthesis/design for three separate objective functions, i.e the optimizations involved the maximization of thrust efficiency, the minimization of fuel mass consumption, and the minimization of exergy destruction plus fuel exergy loss. The principal results found the minimum fuel consumption and minimum exergy destruction measures equivalent, both meeting the constraints of the mission while using 11% less fuel than the thrust efficiency measure. Optimizing the vehicle for the single most constrained mission segment yielded a vehicle capable of flying the entire mission but with fuel consumption and exergy destruction plus fuel loss values greater than the above mentioned integrated vehicle solutions. In essence, the mission-level analysis provided much insight into the dynamics of mission-level hypersonic flight and demonstrated the usefulness of an exergy destruction minimization measure for highly integrated synthesis/design. / Master of Science

This thesis work presents detailed results of the application of exergy-based methods to highly dynamic, integrated aerospace systems such as hypersonic vehicle concepts. In particular, an exergy-based methodology is compared to a more traditional based measure by applying both to the synthesis/design and operational optimization of a hypersonic vehicle configuration comprised of an airframe sub-system and a propulsion sub-system consisting of inlet, combustor, and nozzle components. A number of key design and operational decision variables are identified as those which govern the hypersonic vehicle flow physics and thermodynamics and detailed one-dimensional models of each component and sub-system are developed. Rates of exergy loss as well as exergy destruction resulting from irreversible loss mechanisms are determined in each of the hypersonic vehicle sub-systems and their respective components. Multiple optimizations are performed for both the propulsion sub-system only and for the entire hypersonic vehicle system for single mission segments and for a partial, three-segment mission. Three different objective functions are utilized in these optimizations with the specific goal of comparing exergy methods to a standard vehicle performance measure, namely, the vehicle overall efficiency. Results of these optimizations show that the exergy method presented here performs well when compared to the standard performance measure and, in a number of cases, leads to more optimal syntheses/designs in terms of the fuel mass flow rate required for a given task (e.g., for a fixed-thrust requirement or a given mission). In addition to the various vehicle design optimizations, operational optimizations are conducted to examine the advantage if any of energy exchange to maintain shock-on-lip for both design and off-design conditions. Parametric studies of the hypersonic vehicle sub-systems and components are also conducted and provide further insights into the impacts that the design and operational decision variables and flow properties have on the rates of exergy destruction. / Master of Science

This study examines the effect of varying injectant molecular weight on the penetration of transverse injection jets into a supersonic crossflow. The injectants considered here are methane (W=16.04), air (W=28.97) and carbon dioxide, (W=44.01). These results augment the previous results obtained at Virginia Tech for helium (W=4.00) injection under the same test conditions to provide a very wide range of molecular weights. Second, since shocks are ubiquitous in scramjet combustors, their influence on penetration and mixing was also studied by arranging for an oblique shock to impinge near the injection station. The cases of a shock impinging upstream and downstream of the injector were both examined. One can anticipate an important influence of molecular weight here also because of the importance of density gradients on the generation of vorticity by baroclinic torque. Increasing molecular weight was found to increase penetration in general, as well as increase the lateral spreading of the plume. The majority of the data shows a weak dependency of the jet size on molecular weight, but there are indications that under certain circumstances large changes in the flow structure may occur due to molecular weight effects. The addition of an impinging shock is found to increase mixing and decrease penetration and plume size, especially with the shock impinging downstream of the injector. / Master of Science

The effects of liquid superheat on the disruption of liquid droplets accelerated in a supersonic flow were examined experimentally in a drawdown supersonic wind tunnel. Monodisperse 60 ìm diameter droplets of two test fluids (methanol and ethanol) were generated upstream of the entrance to the tunnel and accelerated with the supersonic flow such that their maximum velocities relative to the air flow were transonic. Droplets were imaged by shadowgraphy and by multiple-exposure direct photography using planar laser sheet illumination. In addition to providing information on droplet lifetime, the latter technique allows measurement of the droplet downstream distance versus time, from which the velocity and acceleration during disruption can be inferred. All droplets were unheated upon injection. Depending on the vapor pressure of the liquid, the droplets achieved varying levels of liquid superheat as they experienced low static pressure in the supersonic flow. Histograms of the droplet population downstream of the supersonic nozzle throat indicate that the lifetime of droplets in supersonic flow decreases with an increasing amount of droplet superheat. The shorter lifetime occurs even as the droplet Weber number (based on initial droplet size) decreases initially due to the lower relative velocity of the methanol droplets to that of ethanol droplets. This is due to a higher acceleration than ethanol droplets of comparable initial size. This is consistent with the more rapid disruption and the faster decrease in mass for the methanol droplets. The droplets, depending on the level of superheating, in some cases underwent disruption modes different than those expected for the corresponding values of Weber number.

This study investigates the design and aeropropulsive performance of a complete, hydrogen powered, shock-induced combustion ramjet (shcramjet) at a flight Mach number of 11 and altitude of 34.5 km. The design includes a Prandtl-Meyer compression inlet, cantilevered ramp fuel injectors, a shock-inducing wedge and a divergent nozzle. Numerical studies are undertaken using the WARP code that solves the three-dimensional Favre-averaged Navier-Stokes equations closed by the Wilcox k-ω turbulence model and the Jachimowski H2/air chemical kinetics model. Studies of fuel injection properties, mixing duct length, combustor wedge and nozzle geometry are completed to maximize the overall performance of the vehicle. The final shcramjet configuration generates a specific impulse of 1110 s. A comparison is undertaken with a scramjet vehicle at identical flight conditions and using many of the same components. The comparable scramjet generates a higher specific impulse of 1450 s although it is significantly larger and therefore heavier.

This study investigates the design and aeropropulsive performance of a complete, hydrogen powered, shock-induced combustion ramjet (shcramjet) at a flight Mach number of 11 and altitude of 34.5 km. The design includes a Prandtl-Meyer compression inlet, cantilevered ramp fuel injectors, a shock-inducing wedge and a divergent nozzle. Numerical studies are undertaken using the WARP code that solves the three-dimensional Favre-averaged Navier-Stokes equations closed by the Wilcox k-ω turbulence model and the Jachimowski H2/air chemical kinetics model. Studies of fuel injection properties, mixing duct length, combustor wedge and nozzle geometry are completed to maximize the overall performance of the vehicle. The final shcramjet configuration generates a specific impulse of 1110 s. A comparison is undertaken with a scramjet vehicle at identical flight conditions and using many of the same components. The comparable scramjet generates a higher specific impulse of 1450 s although it is significantly larger and therefore heavier.

Scramjet engines are the focus of considerable interest for propulsion in the hypersonic flow regime. One of the serious technical challenges for developing scramjets is reducing the skin friction drag on the engine. The combustion chamber, in particular, is a major contributor to the skin friction drag because of the high density of the flow through that region. This investigation focuses on reducing the combustion chamber skin friction drag by minimising the surface area and size of the combustion chamber and by employing a novel approach to accomplishing combustion. The first design criterion is addressed by using a single internal-combustor scramjet configuration, as opposed to multiple external combustors, and by injecting the fuel on the intake to reduce the mixing length required in the combustor. The second design criterion refers to the use of a new technique called radical farming. This uses the highly two-dimensional nature of the flow through the engine, which is created by deliberately ingesting the leading edge shocks, to achieve combustion at lower mean static pressures and temperatures than generally expected. A simplified approximate theoretical analysis of the radical farming concept is presented. Experiments were conducted in the T4 free-piston shock tunnel on a scramjet model with a single rectangular constant cross-sectional area combustion chamber. Pressure measurements were taken along the centreline of the intake, combustion chamber and thrust surface and across the model width at three locations. Gaseous hydrogen fuel was injected halfway along the intake at a range of equivalence ratios between zero and one. The combustion chamber height was varied from 20mm to 32mm, which varied the contraction ratio of the engine from 4.1 to 2.9. The experiments were conducted at a stagnation enthalpy of either 3MJ/kg or 4MJ/kg. The nominal 3MJ/kg condition corresponds to Mach 7.9 flight at an altitude of 24km. The majority of the 4MJ/kg experiments were conducted at a nominal condition corresponding to Mach 9.1 flight at an altitude of 32km. A small number of 4MJ/kg experiments were conducted at simulated flight altitudes of between 30 and 38km; the flight Mach number for these experiments was approximately 9.0. Thrust was calculated by integrating the centreline pressure distribution over the area of the thrust surface, assuming that the pressure at any axial location was constant across the engine width. These experimental thrust values were compared with theoretical estimates obtained using a one-dimensional analysis and a quasi-two-dimensional analysis. The comparison provided an indication of the level of completion of combustion in the experiments. The difference in thrust produced as a result of combusting fuel was examined by plotting the incremental specific impulse against equivalence ratio. Experimental and theoretical results agreed best at the higher equivalence ratios. Turbulent boundary layer separation correlations were used to provide reasonable estimates for the equivalence ratio at which the flow choked. The drag on the internal flowpath of the scramjet engine was estimated using the quasi-two-dimensional analysis. This drag estimate was combined with the experimental thrust measurements to provide estimates of the net specific impulse. Positive net specific impulse estimates were obtained above a certain minimum equivalence ratio, which depended on the contraction ratio and the test condition. The engine performance was observed to be highly dependent on the two-dimensional shock structure within the engine. Thrust and specific impulse were observed to decrease with increasing simulated flight altitude, as expected. Positive net specific impulse estimates were obtained at equivalence ratios of approximately one for simulated flight altitudes below 35km. Assuming complete combustion and that an equivalence ratio of one can be reached, the configuration considered in the present study can theoretically reach a net specific impulse of approximately 1000s at the 3MJ/kg condition and 500s at the 4MJ/kg condition. These numbers provide a promising testimonial for the use of this configuration, with modifications, as a more efficient alternative to rocket engines.

Thesis (Ph.D)--Aerospace Engineering, Georgia Institute of Technology, 2009. / Committee Chair: Menon Suresh; Committee Co-Chair: Jagoda Jeff; Committee Member: Ruffin Stephen; Committee Member: Thorsten Stoesser; Committee Member: Walker Mitchell. Part of the SMARTech Electronic Thesis and Dissertation Collection.

Hypersonic propulsion has become an increasingly important research field over the past fifty years, and subsequent interest in propulsion systems utilizing supersonic combustion has emerged. Air-breathing engines are desirable for such applications as hypersonic flight vehicles would not need to carry an oxidizer. Therefore, hypersonic air-breathing propulsion systems require an inlet with high mass capture and compressive efficiency. The present work seeks to outline the development and validation of a novel design tool for producing air inlet designs for hypersonic vehicles at variable flight conditions. A Busemann inlet was chosen for its high compressive efficiency, geometric flexibility, and existing experimental validation. The design tool uses the Taylor-Maccoll equation to generate a streamline through a conical flow field. A streamline tracing technique is used to produce three-dimensional inlet surfaces with various capture areas. Additionally, a surface morphing process is implemented to combine inlet profiles for improved engine compatibility. The inlet morphing process allowed for the creation of inlets with offset exit profiles. These offset profiles were evaluated at off-design Mach numbers using Star-CCM+ to quantify efficiency metrics and characterize starting phenomena.