No description available.

The subject investigation demonstrated that it is feasible to use the microwave technique to measure radial burning rates ,and differential burning rates in solid propellant rocket motors. A simulator, consisting of a spiral rotating in an oil bath, was used to represent the curved burning surface of a tubular grain of propellant with the outer surface and ends restricted. The radial movement of the spiral, simulating radial burning, was detected by recording the phasor difference of the reflected microwaves from the reflecting surface of the spiral and the reflected microwaves from a stationary reference surface. The reflected microwaves changed in phase relation producing successive minimum values in the detected signal for each one-half of a microwave wavelength in oil displacement of the spiral. The displacement rates were calculated as average rates for a displacement of one-half of a microwave wavelength in oil. The curved reflective surface did not present a measurement problem. The differential displacement rates were detected by recording the phasor difference of the reflected signals from two spirals. The reflected signals changed in phase relation, if the reflecting surfaces were moving at different rates, producing a beat frequency in the detected signal. The differential displacement rate was determined from the number of beat frequency cycles, the one-half microwave wavelength in oil, and the time. The addition of the aluminum powder to the oil simulating aluminized propellants did not prevent detection of the moving surface. The results indicated that the microwave technique can be applied to aluminized propellants. / M.S.

Polybutadienes, polymerized via free radical mechanism to give an average molecular weight of 3000, were analyzed with carbon-13 NMR. The relative abundance of the three types of structural units (cis-1,4, trans-1,4, and vinylic-1,2 units) was quantitatively determined. The distribution of the three structural units was found to be completely random. Branching in the analyzed polymer was determined to be low (estimated to be less than 3 %). In hydroxyl-terminated-polybutadiene, two separate resonance signals from the hydroxyl-bearing end carbons were observed. Six isomeric dinitrotoluenes and four isomeric trinitrotoluenes were characterized with carbon-13 chemical shifts, which can be used in the qualitative and quantitative analysis of mixtures of these compounds. Multiple linear regression analysis was performed on the chemical shifts to obtain parameters which are useful in estimating the chemical shifts of carbon-13 nuclei of methyl- and nitro- substituted benzenes. Carbon-13 chemical shifts of other propellant ingredients (aliphatic nitrate esters, carboranes, plasticizers and stabilizers) are reported. / Master of Science

A flow visualization study was conducted to analyze flow characteristics inside the solid rocket motor (SRM) used on the NASA space shuttle. The objective of this investigation was to determine whether the internal flow structure could adversely affect the nozzle/case joint and the surrounding casing. Also, it was hoped to learn more about causes of low level acoustic pressure oscillations observed during SRM test firings. The SRM was simulated by water flow through a plexiglas model mounted in a water tunnel. Dye and hydrogen bubble visualization techniques along with hot water analysis methods were used to detect flow patterns. Visual results recorded on video tape indicated strong circumferential and recirculation flows around the nozzle. Vortex formation near the nozzle inlet was also observed and was the prime focus of this investigation. Because the nozzle inlet geometry was very similar to an aircraft engine inlet operating close to the ground, vortices seen in this investigation were believed to behave like vortices seen around engine inlets. Based on the results from this investigation and the results of previous engine inlet vortex studies, it was concluded that the nozzle vortices could be the excitation source of SRM pressure oscillations. / Master of Science

No description available.

This work describes the development of a hybrid rocket propulsion system for a reusable sounding rocket, as part of the first phase of the UKZN Phoenix Hybrid Sounding Rocket Programme. The programme objective is to produce a series of low-to-medium altitude sounding rockets to cater for the needs of the African scientific community and local universities, starting with the 10 km apogee Phoenix-1A vehicle. In particular, this dissertation details the development of the Hybrid Rocket Performance Code (HRPC) together with the design, manufacture and testing of Phoenix-1A’s propulsion system. The Phoenix-1A hybrid propulsion system, generally referred to as the hybrid rocket motor (HRM), utilises SASOL 0907 paraffin wax and nitrous oxide as the solid fuel and liquid oxidiser, respectively. The HRPC software tool is based upon a one-dimensional, unsteady flow mathematical model, and is capable of analysing the combustion of a number of propellant combinations to predict overall hybrid rocket motor performance. The code is based on a two-phase (liquid oxidiser and solid fuel) numerical solution and was programmed in MATLAB. HRPC links with the NASA-CEA equilibrium chemistry programme to determine the thermodynamic properties of the combustion products necessary for solving the governing ordinary differential equations, which are derived from first principle gas dynamics. The combustion modelling is coupled to a nitrous oxide tank pressurization and blowdown model obtained from literature to provide a realistic decay in motor performance with burn time. HRPC has been validated against experimental data obtained during hot-fire testing of a laboratory-scale hybrid rocket motor, in addition to predictions made by reported performance modelling data. Development of the Phoenix-1A propulsion system consisted of the manufacture of the solid fuel grain and incorporated finite element and computational fluid dynamics analyses of various components of the system. A novel casting method for the fabrication of the system’s cylindrical single-port paraffin fuel grain is described. Detailed finite element analyses were performed on the combustion chamber casing, injector bulkhead and nozzle retainer to verify structural integrity under worst case loading conditions. In addition, thermal and pressure loading distributions on the motor’s nozzle and its subsequent response were estimated by conducting fluid-structure interaction analyses. A targeted total impulse of 75 kNs for the Phoenix-1A motor was obtained through iterative implementation of the HRPC application. This yielded an optimised propulsion system configuration and motor thrust curve. / Thesis (M.Sc.Eng.)-University of KwaZulu-Natal, Durban, 2013.

Basic experimental studies of the injection of liquid fuel into a two dimensional flowfield designed to represent a sudden-expansion "dump" combustor were performed under cold-flow conditions. Test conditions were as follows: 0.6 entrance Mach number, 25 PSIA total pressure, and nominally 75°F stagnation temperature. Two step heights were investigated, 1.0 in. and 0.5 in., corresponding to area ratios of 1.33 and 1.17. The investigation included Pitot and static pressure distributions, spark and streak shadowgraphs, surface flow visualization, direct photographs and videotape recordings. The backlighted streak and spark shadowgraphs were used to obtain jet penetration and break-up information. Oil drop surface flow studies showed details of the flow in the recirculation region behind the step. The injectant for these cold flow studies was selected as water, which was injected transversely to the air flow 1.0 in. and 0.5 in. upstream of the step at various flow rates. It was found that both the location of the injection port relative to the step and the step height had no measurable effect on jet penetration and break-up. Injectant accumulation on the combustor wall in the base-flow region was found to be substantial under some conditions, and the amount of accumulation was shown to be a strong function of initial liquid jet penetration height. / M.S.

A dual-frequency microwave bum rate measurement system for solid rocket motors has been developed and is described. The system operates in the X-band (8.2-12.4 Ghz) and uses two independent frequencies operating simultaneously to measure the instantaneous bum rate in a solid rocket motor. Modeling of the two frequency system was performed to determine its effectiveness in limiting errors caused by secondary reflections and errors in the estimates of certain material properties, particularly the microwave wavelength in the propellant. Computer simulations based upon the modeling were performed and are presented. Limited laboratory testing of the system was also conducted to determine its ability perform as modeled. Simulations showed that the frequency ratio and the initial motor geometry (propellant thickness and combustion chamber diameter) determined the effectiveness of the system in reducing secondary reflections. Results presented show that higher frequency ratios provided better error reduction. Overall, the simulations showed that a dual frequency system can provide up to a 75% reduction in burn rate error over that returned by a single frequency system. The hardware and software for dual frequency measurements was developed and tested, however, further instrumentation work is required to increase the rate at which data is acquired using the methods presented here. The system presents some advantages over the single frequency method but further work needs to be done to realize its full potential. / Master of Science

Call number: LD2668 .T4 1963 R93 / Master of Science

A method for developing an erosive burning model for use in solid propellant design-and-analysis interior ballistics codes is described and evaluated. Using Direct Numerical Simulation, the primary mechanisms controlling erosive burning (turbulent heat transfer, and finite rate reactions) have been studied independently through the development of models using finite rate chemistry, and infinite rate chemistry. Both approaches are calibrated to strand burn rate data by modeling the propellant burning in an environment with no cross-flow, and adjusting thermophysical properties until the predicted regression rate matches test data. Subsequent runs are conducted where the cross-flow is increased from M=0.0 up to M=0.8. The resulting relationship of burn rate increase versus Mach Number is used in an interior ballistics analysis to compute the chamber pressure of an existing solid rocket motor. The resulting predictions are compared to static test data. Both the infinite rate model and the finite rate model show good agreement when compared to test data. The propellant considered is an AP/HTPB with an average AP particle size of 37 microns. The finite rate model shows that as the cross-flow increases, near wall vorticity increases due to the lifting of the boundary caused by the side injection of gases from the burning propellant surface. The point of maximum vorticity corresponds to the outer edge of the APd-binder flame. As the cross-flow increases, the APd-binder flame thickness becomes thinner; however, the point of highest reaction rate moves only slightly closer to the propellant surface. As such, the net increase of heat transfer to the propellant surface due to finite rate chemistry affects is small. This leads to the conclusion that augmentation of thermal transport properties and the resulting heat transfer increase due to turbulence dominates over combustion chemistry in the erosive burning problem. This conclusion is advantageous in the development of future models that can be calibrated to heat transfer conditions without the necessity for finite rate chemistry. These results are considered applicable for propellants with small, evenly distributed AP particles where the assumption of premixed APd-binder gases is reasonable.