Robert Stowe

Polymer-Grafted Metal Nanoparticles for Fuel Applications

Ultrafine metal powders have been identified as very promising fuels for future energetic material formulations. However, the large specific surface area that gives these powders a high reactivity also makes them particularly difficult to remain in a nonoxidized state. They also agglomerate easily during compounding processes due to strong particle-to-particle interactions. The coating of the particles with a polymer may offer a solution to these problems. We investigated two in situ polymerization processes using thermoplastic and thermoset coatings. Polyolefins such as polyethylene and polypropylene were obtained using a modified Ziegler-Natta reaction scheme. This process was found to be flexible enough to control the amount of polyethylene grafted onto the powders. The second type of coating was based on polyurethane chemistry. Nanometric-sized aluminum and boron powders were treated and characterized by means of thermogravimetric analysis, electronic microscopy, and x-ray photoelectron microscopy. The barrier properties of the polymer layer grafted onto the particles were evaluated using a chemical digestion method and thermoanalytical techniques. Polyethylene-coated particles showed a better resistance to early aging under stringent conditions of humidity and temperature and therefore would be expected to demonstrate a longer shelf life in a propellant formulation.

Design Methodology for a Pulse Detonation Engine as a Ramjet Replacement

An idealized pulse detonation engine (PDE) based on a realistic system concept is defined for which propulsive performance is significantly greater than that of a ramjet over a large range of Mach numbers. An analytical model of the system is developed that calculates physical, timing and flow parameters that can be used as input to multi -dimensional computational fluid dynamics (CFD) analyses and provide a starting point for real system design. A parametric study of the influence on propulsive performance of flight Mach number, inlet pressure recovery, l ength of the detonation tube, mass flow rate of air through the engine, percentage purge and the tube filling strategy is performed with an analytical model and a one -dimensional code based on the method of characteristics. Finally, in a sample engine des ign exercise a PDE based system is specified for which propulsive performance for steady flight is significantly greater than that of a ramjet .

Formation and Structure of Steady Oblique and Conical Detonation Waves

The formation and structure of oblique detonation waves initiated by semi-infinite wedges and cones are presented. For wedge or cone angles less than the deflection angle required for an oblique Chapman–Jouguet (CJ) detonation, different wave structures have been previously reported. Using the method of characteristics and numerical simulations, it is shown that, for such low wedge or cone angles, a CJ oblique detonation is eventually initiated following an induction process. It its thus demonstrated that shock-induced combustion with the reaction front remaining uncoupled to the oblique shock in the far field is not a valid solution. Simulations with semi-infinite cones reveal that the effect of the front curvature around the cone axis allows oblique detonations to be formed at angles lower than that of a planar CJ oblique detonation.

Performance Prediction of a Ducted Rocket Combustor Using a Simulated Solid Fuel

The ducted rocket is a supersonic flight propulsion system that takes the exhaust from a solid fuel gas generator, mixes it with air, and burns it to produce thrust. To develop such systems, the use of numerical models based on computational fluid dynamics (CFD) has been increasing, but to date only simplified treatments of the combustion within ducted rockets have been reported, likely due to the difficulties in characterizing and accurately modeling the partially reacted, particle-laden fuel exhaust from the gas generator. Through a careful examination of the governing equations and experimental measurements, a CFD-based methodology that properly accounts for the influence of the gas generator exhaust, particularly the solid phase, has now been developed to predict the performance of a ducted rocket combustor using a simulated solid fuel. It uses an equilibrium-chemistry probability density function combustion model with two separate streams, one gaseous and the other of 75-nm-diam carbon spheres, to represent the exhaust products from the gas generator. After extensive validation with direct-connect combustion experiments over a wide range of geometries and test conditions, this CFD-based method was able to predict, within a good degree of accuracy, the combustion efficiency of a ducted rocket combustor.

Deuterium Effect on Thermal Decomposition of Deuterated Gap: 1. Slow Thermal Analysis with a TGA-DTA-FTIR-MS

The knowledge of the degradation mechanisms of propellants is a key factor in understanding some performance issues of chemical propulsion systems. The thermal decomposition of GAP and a deuterated analog was studied. Thermogravimetric analyses coupled to differential thermal analysis, Fourier transform infrared spectrometry, and mass spectrometry were employed to evaluate the complete decomposition profile of GAPs. Evidence of labeled gaseous compounds were found both in mass spectra and in the FTIR spectra, where new frequencies associated with isotopic labeled gaseous products were detected. An imine-bearing species was detected among the decomposition products, apparently originating from the breakdown of the azide group.

Generalized Formulation for Droplet Size Distribution in a Spray

Based on experimental data in the open literature, a generalized formulation is proposed for droplet size distribution in liquid sprays. The methodology is based on the use of the Pearson system of distribution curves. Experimental data demonstrated that the beta family (type I) distribution in the Pearson system was able to fully characterize the whole data. It was also verified that the gamma family (type III) could be used as a simplified model with a comparable accuracy to a Nukiyama-Tanasawa distribution or better than a Rosin-Rammler.

Modelling Combustor Performance of a Ducted Rocket

Under a co-operative program, the Defence Research Establishment Valcartier and Universite Laval in Canada and the TNO Prins Maurits Laboratory in the Netherlands have studied the use of a ducted rocket for missile propulsion. Previously, non-reacting flow modelling using CFD (Computational Fluid Dynamics) and water tunnel tests were performed on a wide range of configurations to identify the geometries and flow rates necessary for good air/fuel mixing in the ramjet combustor. Hot-flow direct-connect combustion experiments using both simulated and solid fuels have since been carried out on these same configurations to measure actual combustor performance. Using a PDF (Probability Density Function) combustion model, reacting flow CFD modelling has also been done for each of these configurations with the goal of being able to analyze and predict combustor performance. Agreement between the measured and calculated temperature-based efficiencies was good for some configurations, but the overall tendency was to overestimate with the CFD. The differences are likely due to incorrectly specifying the boundary conditions and inaccuracies due to how the flowfield, turbulence and combustion were modelled.

Turbulent Flow Effects on DDT Run-up Distance for a Pulse Detonation Engine

Experiments were performed to investigate the effect of increased flow rate and turbulence on the deflagration-to-detonation transition (DDT) run-up distance of a propane/oxygen/nitrogen gas mixture within a smooth bore tube. The mixture was introduced through the intake valves of an automobile cylinder head into the tube and varied from 588 l/min to 3800 l/min. Turbulence was introduced into the flowing mixture by the motion of the intake valves, which generated vortex shedding and swirling. The results indicated that the DDT run-up distance was decreased considerably, by as much as 6 times that of quiescent flow, as the flow Reynolds number was increased. Results from gas mixture for various β-values of 0.0, 0.46 and 0.64 showed similar trends for the reduction of DDT run-up distance with increased Reynolds number. A minimum value of DDT distance for a unique value of Reynolds number was also observed. Increasing the cyclic operation of the valves from 5Hz to 20Hz for a given flow Reynolds number also resulted in a decrease of DDT run-up distance for each of the flow rates tested, further implying that increased turbulence decreased the DDT run-up distance.

TRANSIENT CHAMBER FLOWFIELD SIMULTATION OF A ROD-AND-TUBE CONFIGURATION SOLID ROCKET MOTOR

A transient CFD simulation of the flowfield within a rod-and-tube solid propellant rocket motor has been developed. This model couples the fluid dynamics and heat transfer of the gas flowfield within the rocket port to the nozzle to predict the internal environment within the motor including the regression rate of the propellant. The propellant regression is described with an empirical erosive burning model based on the phenomenological heat transfer approach derived by Lenoir and Robillard 1 . The predicted propellant burn rate and consequently the chamber pressure were found to be significantly increased from the case where propellant regression was described by the simple burning law only. This augmentation of the burn rate, particularly during the early stages of the simulation, was in agreement with the trends observed in small diameter rockets where erosive burning was present. A validation of the model comparing an actual Pressure - Time plot to that predicted by the CFD model was also carried out and achieved a very high degree of correlation.

TWO PHASE FLOW COMBUSTION MODELLING OF A DUCTED ROCKET

Under a co-operative program, the Defence Research Establishment Valcartier and Universite Laval in Canada and the TNO Prins Maurits Laboratory in the Netherlands have studied the use of a ducted rocket for missile propulsion. Hot-flow direct-connect combustion experiments using both simulated and solid fuels have been carried out on a wide range of configurations to identify the geometries and flow rates necessary for good combustor performance. The experiments using a simulated ducted rocket fuel, a reacted mixture of ethylene and air, have all been modelled using reacting flow Computational Fluid Dynamics (CFD) with the goal of being able to analyze and predict combustor performance. The combustion was modelled with onestream and twostream PDF (Probability Density Function) models. With the onestream model, all of the fuel components, both gaseous and solid carbon, were injected together and were assumed to react instantaneously in the presence of the oxidizer. Because of this, the onestream model overpredicted the combustion efficiency with respect to the experimental results for most of the combustor configurations examined. With the twostream model, however, the fuel stream was separated into gaseous and solid carbon components, with the carbon injected as a series of 75 nm particles. These particles decompose gradually into carbon monoxide gas, based on a model using both the kinetics of the surface reactions and the diffusion of oxygen to the surface of the particles. For the majority of the configurations, better predictions of combustion efficiency were obtained with the twostream approach when compared to the experimental results than for the onestream PDF model.