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.

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.

Supersonic inlet starting by mass spillage is studied theoretically and numerically in the present thesis. A quasi-one-dimensional, quasi-steady theory is developed for the analysis of flow inside a perforated inlet. The theory results in closed-form relations applicable to flow starting by the mass spillage technique in supersonic and hypersonic inlets. / The theory involves three parameters to incorporate the multi-dimensional nature of mass spillage through a wall perforation. Mass spillage through an individual slot is studied to determine these parameters; analytical expressions for these parameters are derived for both subsonic and supersonic flow conditions. In the case of mass spillage from supersonic flows, the relations are exact. However, due to the complexity of flow field, the theory is an approximation for subsonic flows. Therefore, a correction factor is introduced which is determined from an empirical relation obtained from numerical simulations. / A methodology is also proposed to determine perforation size and distribution to achieve flow starting for a given inlet at a desired free-stream Mach number. The problem of shock stability inside a perforated inlet designed with the proposed method is also discussed. / The method is demonstrated for some test cases. Time-realistic CFD simulations and experimental results in the literature confirm the accuracy of the theory and the reliability of the proposed design methodology.

No description available.

In the pursuit of developing hypersonic cruise vehicles, unstart is a major roadblock to achieving stable flight. Unstart occurs when a sudden instability in the combustor of a vehicle’s propulsion system creates an instantaneous pressure rise that initiates a shock. This shock travels upstream out of the inlet of the vehicle, until it is ejected from the inlet and creates a standing shockwave that chokes the flow entering the vehicle, thereby greatly reducing its propulsive capability. In severe cases, this can lead to the loss of the vehicle. This thesis presents the results of a computational study of the dynamics of unstart near Mach 5 and presents some possible precursor signals that may indicate its presence in flight. Using SU2, an open-source CFD code developed at Stanford University, the Unsteady Reynolds-Averaged Navier-Stokes equations are used to develop a model for flow in a scramjet inlet-isolator geometry, both in the fully started state and during unstart. The results of these calculations were compared against experimental data collected by J. Wagner, at the University of Texas, Austin. In the present computations, unstart was initiated through the use of an artificial body force, which mimicked a moveable flap used in the experiments. Once the results of the code were validated against these experiments, a selection of parametric studies were conducted to determine how the design of the inlet-isolator by Wagner affected the flow, and thus how generalizable the results can be. In addition, precursor signals indicative of unstart were identified for further study and examined in the different parametric studies. It was found that a thick boundary layer is conducive to a stronger precursor signal and a slower unstart. In addition, an aspect ratio closer to 1:1 promotes flow mixing and reduces the unstart speed and strength. Moreover, an aspect ratio in this range reduces the precursor signal strength but, if a thick boundary layer is present, will smear the signal out over a larger area, potentially making it easier to detect. <br>

The purpose of this research was to determine if it is feasible to apply Lorentz mixing to supersonic diffusion flames, such as those found in SCRAMjet engines. The combustion rate in supersonic diffusion flames is limited by the rate at which air and fuel mix. Lorentz mixing increases turbulence within a flow, which increases the rate at which species mix and thus increases the rate of combustion. In order to determine the feasibility of Lorentz mixing for this application, a two-dimensional model of supersonic reacting flow with the application of a Lorentz force has been examined numerically. The flow model includes the complete Navier-Stokes equations, the ideal gas law, and terms to account for diffusion of chemical species, heat release due to chemical reaction, change in species density due to chemical reaction, and the Lorentz forces applied during Lorentz mixing. In addition, the Baldwin-Lomax turbulence model is used to approximate turbulent transport properties. A FORTRAN program using the MacCormack method, a commonly used computational fluid dynamics algorithm, was used to solve the governing equations. The accuracy of the program was verified by using the program to model flows with known solutions. Results were obtained for flows with Lorentz forces applied over a series of power levels and frequencies. The results show significant increases in the rate of combustion when Lorentz mixing is applied. The amount of power required to drive Lorentz mixing is small relative to the rate at which energy is released in the chemical reaction. An optimum frequency at which to apply Lorentz mixing was also found for the flow being considered. The results of the current study show that Lorentz mixing looks promising for increasing combustion rates in supersonic reacting flows, and that future study is warranted. In particular, researchers attempting to improve combustion in SCRAMjet engines may want to consider Lorentz mixing as a way to improve combustion. / Graduation date: 1997

The main objective of this thesis is to develop a comprehensive model with the capability of modeling both a high Reynolds number and high magnetic Reynolds number turbulent flow for application to supersonic combustor. The development of this model can be divided into three categories: one, the development of a self-consistent MHD numerical model capable of modeling magnetic turbulence in high magnetic Reynolds number applications. Second, the development of a gas discharge model which models the interaction of externally applied fields in conductive medium. Third, the development of models necessary for studying supersonic combustion applications with plasma-assistance such the extension of chemical kinetics models to extremely high temperature and non-equilibrium phenomenon.

One of the current goals of research in hypersonic, airbreathing propulsion is access to higher Mach numbers. A strong driver of this goal is the desire to integrate a scramjet engine into a transatmospheric vehicle airframe in order to improve performance to low Earth orbit (LEO) or the performance of a semi-global transport. An engine concept designed to access hypervelocity speeds in excess of Mach 10 is the shock-induced combustion ramjet (i.e. shcramjet). This dissertation presents numerical studies simulating the physics of a shcramjet vehicle traveling at hypervelocity speeds with the goal of understanding the physics of fuel injection, wall autoignition mitigation, and combustion instability in this flow regime. This research presents several unique contributions to the literature. First, different classes of injection are compared at the same flow conditions to evaluate their suitability for forebody injection. A novel comparison methodology is presented that allows for a technically defensible means of identifying outperforming concepts. Second, potential wall cooling schemes are identified and simulated in a parametric manner in order to identify promising autoignition mitigation methods. Finally, the presence of instabilities in the shock-induced combustion zone of the flowpath are assessed and the analysis of fundamental physics of blunt-body premixed, shock-induced combustion is accelerated through the reformulation of the Navier Stokes equations into a rapid analysis framework. The usefulness of such a framework for conducting parametric studies is demonstrated.

No description available.

The recent resurgence in hypersonics is centered around the development of SCRAMJET engine technology to power future hypersonic vehicles. Successful flight trials by Australian and American scientists have created interest in the scramjet engine research across the globe. To develop scramjet engine, it is important to study heat transfer effects on the engine performance and aerodynamic forces acting on the body. Hence, the main aim of present investigation is the design of scramjet engine configuration and measurement of aerodynamic forces acting on the model and heat transfer rates along the length of the combustor. The model is a two-dimensional single ramp model and is designed based on shock-on-lip (SOL) condition. Experiments are performed in IISc hypersonic shock tunnel HST2 at two different Mach numbers of 8 and 7 for different angles of attack. Aerodynamic forces measurements using three-component accelerometer force balance and heat transfer rates measurements using platinum thin film sensors deposited on Macor substrate are some of the shock tunnel flow diagnostics that have been used in this study.