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.

Thesis (M.App.Sc.) - Carleton University, 2007. / Includes bibliographical references (p. 194-202). Also available in electronic format on the Internet.

Cette thèse présente une étude des instabilités haute-fréquence dans les moteurs-fusées. Ce phénomène, qui a posé de nombreux problèmes dans les programmes de développement de moteur, est abordé de trois façons complémentaires : expérimentalement, théoriquement et numériquement. Premièrement, des expériences sont menées afin d’identifier les principaux processus et d’apporter les mécanismes ayant lieu lorsque le moteur devient instable. Pour parvenir à ce stade, un nouveau modulateur (VHAM), capable de créer des ondes acoustiques représentatives de ce qui se produit dans un moteur réel, est conçu et caractérisé. La deuxième partie concerne l’analyse théorique. Deux modèles (FAME, SDM) sont développés en suivant les principales conclusions de la campagne expérimentale : les oscillations de dégagement de chaleur sont dues au mouvement transverse des flammes, et le phénomène est déclenché lorsque des gouttelettes deviennent suffisamment petites pour être convectées par le champ acoustique. En utilisant ces modèles comme base de référence, un code numérique (STAHF) est présenté. Son but est de rendre compte des mécanismes déjà identifiés pour un coût de calcul faible. Il est ensuite montré qu’il peut être utilisé pour étudier des moteurs-fusées grandeur nature. La LES compressible est choisie pour étudier l’interaction entre l’acoustique et la combustion numériquement. Un nouveau modèle de combustion pour flammes non-prémélangées basé sur une hypothèse de chimie infiniment rapide est présenté et validé sur une flamme bien documentée (H3). Il est ensuite utilisé pour étudier l’interaction entre une onde acoustique transverse et la flamme H3. Une comparaison entre le terme source de Rayleigh calculé à partir de la simulation et celui prédit par le modèle théorique FAME est finalement menée. / This thesis presents a study of high frequency instabilities in rocket engines. This issue, which has plagued many engine development programs, is approached by three complementary viewpoints: experimental, theoretical, and numerical. First, experiments are carried out to identify the main processes involved and bring forth mechanisms taking place when an engine becomes unstable. To achieve this stage, a new modulator (the VHAM), capable of creating acoustic waves representative of what occurs in an actual engine, is designed and characterized. The second part of this thesis concern theoretical analysis. Two models are developed following the main conclusions of the experimental campaign: heat release oscillations are due to the transverse flames’ motion, and the phenomenon is triggered when droplets become small enough to be convected by the acoustic field. Using these models as a baseline, a numerical code (STAHF) is presented. Its purpose is to account for mechanisms identified previously for little computational cost. This code is validated on particularly responding situations observed during experiments. It is then shown that it can be used to study real scale rocket engines. The third point of view adopted to address the problem is numerical simulation. Full compressible LES is chosen to study the interaction between acoustics and combustion. A new combustion model for non-premixed flames with infinitely fast chemistry is presented and validated on a well documented flame (H3). It is then used to study the interaction between a transverse acoustic wave and the H3 flame. A comparison between the Rayleigh source term computed from the simulation and the one predicted by the theoretical model FAME is conducted eventually.

Energetic materials (i.e., explosives, propellants, and pyrotechnics) have been used for centuries in a wide variety of applications that include celebratory firework displays, the demolition of ‘immovable’ structures, mining resources from the earth’s crust, launching humans into outer space, and propelling munitions across the battlefield. Many different scientific and engineering domains have found unique value in their characteristic release of significant heat and pressure. While the rate at which energetic materials react is often dependent on the source of initiation, surrounding thermodynamic conditions, and formulation sensitivity, many applications aim for a controlled combustion process to produce large amounts of work output – solid and liquid rocket motors and gun-launched projectiles are a few key examples. Other energetic material systems are often inadvertently exposed to thermal insults, which can result in similar combustion behavior. To accurately model these systems, it is important to have a fundamental understanding of the chemical kinetics that control various aspects of the combustion process (e.g., changes in temperature (T), pressure (P), and species mole fractions (X)). Detailed chemical kinetic models are often used to understand and subsequently predict such behavior. Understanding the gas-phase reaction kinetics of energetic materials is essential when trying to predict critical performance parameters such as flame speeds, temperature and pressure profiles, and heat flux between material phases. These parameters can have significant impact on predictions of system-level performance (e.g., the specific impulse of solid rocket motors, propellant burn rates in projectile systems, and munition responses to thermal insult and extended temperature cycling). While the gas-phase reaction kinetics of energetic material combustion were heavily studied from the late 1970’s to the early 2000’s, research efforts beyond this time frame have primarily focused on condensed-phase chemistry as it is thought to be less understood. Over the past two decades, however, there have been significant advances in our understanding of small molecule reactions that have not yet been accounted for in many energetic material models. One such example are chemically termolecular reactions – a new class of phenomenological reactions that have not yet been considered for inclusion in any energetic material kinetic models. Recent studies have indicated that chemically termolecular reactions, mediated through ephemeral collision complexes, have significant impact on the global kinetics of certain combustion systems. This discovery has since prompted the question of which systems are significantly influenced by chemically termolecular reactions and should therefore account for their presence in gas-phase phenomenological models. Although a select number of systems have already been investigated, such as flame speed and ignition delay predictions in common hydrocarbon combustion scenarios, the influence of chemically termolecular reactions on the kinetics of energetic materials has not yet been explored. As an initial investigation into energetic materials, a case study for RDX was performed, for which abundant computational and experimental data are available. To aid in assessing the impact of chemically termolecular reactions, for which almost no data are available, this study leveraged an automated procedure to identify and estimate rate constants for potential chemically termolecular reactions based exclusively on data available for related reactions. Four detailed kinetics models for RDX were independently screened for potential chemically termolecular reactions. Model predictions including these chemically termolecular reactions revealed that they have significant potential impact on profiles of major species, radicals, and temperatures. T he analysis pinpointed ∼20-40 chemically termolecular reactions, out of the thousands of possibilities, estimated to have the largest impact. These reactions, including many mediated by ephemeral HNO** and NNH** complexes, are therefore worthwhile candidates for more accurate quantification via master equation calculations. More generally, just as the importance of including chemically termolecular reactions in hydrocarbon combustion models is becoming recognized, the present results show compelling evidence for the need for their inclusion in energetic material models as well. The investigation into chemically termolecular reactions yielded a secondary conclusion based on the observed influence of the small molecule C/H/N/O chemistry on overall predictions of energetic material combustion – updating the small molecule chemistry in RDX models produced significant changes to predictions of major species and temperature, suggesting that the development of a comprehensive gas-phase energetic material combustion model would be of great value and have broad utility as a foundational model for a great variety of C/H/N/O energetic materials. To begin developing such a model, all small molecule chemistry in current kinetic models was reviewed with the intent of identifying a sub-model in need of revisions and subsequently addressing its uncertainties using targeted experiments to improve overall predictions. The ammonia sub-model was selected as it is both highly uncertain and highly influential in many energetic material models. Ammonia (NH₃) has garnered substantial attention in recent years due to its importance across many scientific domains – including its potential use as a carbon-free fuel and long-term energy storage option, its use in reducing combustion-generated nitrogen oxide emissions, its role as a decomposition fragment of many energetic materials, and its presence as an important impurity during biofuel and biomass combustion that can affect overall system kinetics, among others. Yet, it is generally recognized that there are still significant gaps in the present understanding of ammonia kinetics -– in both experimental data sets and sub-models within the overall ammonia kinetic mechanism. For example, most experimental studies of ammonia oxidation have used molecular oxygen as the primary or sole oxidizer. While large mole fractions of molecular oxygen are encountered in many combustion scenarios, there are select systems where ammonia is more likely to be oxidized via nitrogen-containing species (e.g. N₂O and NO₂) and, more generally, there are relatively untested reaction sets that would be accentuated in such conditions. To address these gaps in available experimental data needed for the validation of ammonia kinetics models, jet-stirred reactor experiments were performed for mixtures of NH₃/N₂O/N₂ over an intermediate temperature range (850-1180 K). In these experiments, the mole fractions of NH₃, N₂O, and NO were measured using a combination of gas chromatography, chemiluminescence, electrochemical detection, and infrared absorption – where agreement among the different diagnostics (within 3% for N₂O and 7% for NO) ensured high confidence in the experimental measurements. Comparison of the experimental results and model predictions suggested deficiencies in commonly used models for nitrogen kinetics. Various modeling analyses pointed to the central role of the N₂O + NH₂ = N₂H₂ + NO reaction, on which recent kinetic models all rely on the same rate constant estimate that appears to have not been tested in previous validation data sets for NH₃ kinetics. A second set of jet-stirred reactor experiments were performed for mixtures of NH₃/NO₂/O₂/N₂ over a slightly different temperature range (700–1100 K). Agreement among different diagnostics (≤7% for NO₂ and ≤4% for NH₃) and excellent experimental repeatability confirmed high confidence in all species measurements. Measured mole fractions were compared to predictions from five recently developed kinetic models using flux analysis and uncertainty-weighted kinetic sensitivity analysis, both of which pointed to the importance of reactions involving H₂NO that are both influential in this system and highly uncertain. The measurements from the jet-stirred reactor experiments presented here were combined with comprehensive sets of experimental data and high-level theoretical kinetics calculations using the MultiScale Informatics (MSI) approach to unravel the large uncertainties present in current NH3 oxidation kinetic sub-models. Emphasis was placed on NH₃ oxidation via nitrogen-containing species as this chemistry has been shown to accentuate influential reactions (e.g., the NO₂+NH₂ and NH₂+NO reactions) that are known to be important during the combustion of many energetic materials (e.g., AN, ADN, and AP). The resulting MSI model accurately predicted nearly all of the experimental and theoretical target data within estimated or reported uncertainties. Additional predictions of two NH₃/NO₂ validation data sets, which were not included in the MSI framework, demonstrated its ability to accurately extrapolate predictions to untested T/P/X conditions, indicating that the converged MSI model demonstrates truly predictive behavior. The MSI NH₃ oxidation model presented here should be considered for inclusion in many energetic material models as the NH₃/NOₓ kinetic system is known to be important to the combustion of various propellant and explosive formulations. This sub-model will help to form a foundational gas-phase kinetic model relevant to many different energetic materials, including those that contain inorganic additives for increased energy density and blast effects.

Using the Debye-Huckle approximation, the effects of Coulomb interactions on the equilibrium, frozen, and nonequilibrium flow of an ionized gas have been investigated. The gas is assumed to be monatomic, electrically neutral, and thermal equilibrium (i.e., a one temperature fluid); but the composition of the gas is arbitrary, that is, multiple ionization of any degree is allowed. The thermodynamic variables are derived starting from the appropriate expression for the Helmholtz free energy. Using Boltzmann statistics and assuming that the velocity distribution functions are given by their Maxwellian values, the rate of ionization is derived for atom-atom, atom-ion, and atom-electron collisions. The resulting expressions are then employed in solving the quasi-one-dimensional flow in a converging-diverging nozzle for the equilibrium, frozen, and nonequilibrium cases. Numerical examples, using argon as the working substance, are discussed and the results presented graphically. The results of these calculations indicate that, for single ionization, the effect of Coulomb interactions on the performance of rocket engines is negligible; but that data obtained from hypersonic arc jet wind-tunnels can be significantly influenced by the presence of the interactions. / Ph. D.

The analysis and redesign of a centrifugal impeller for a rocket pump is presented in this thesis. A baseline impeller was designed by Rocketdyne for the NASA Marshall Pump Consortium. Initially, the objective was to reduce the circumferential exit flow distortion of the baseline impeller. Later in the study, the objective became raising the head coefficient of the impeller. The study presented in this thesis was also undertaken to demonstrate current CFD capabilities for impeller design. A literature review includes an overview of centrifugal impeller geometries and configurations. Centrifugal impeller performance and secondary flows are discussed, and a summary of studies on the effects of impeller exit and diffuser inlet velocity distortion on diffuser performance is also presented. The flow calculation details and the results of the baseline impeller flow calculations are described. Fourteen redesigned impeller geometries were analyzed using the Moore Elliptic Flow Program, and the results were compared to the baseline geometry in terms of head rise, losses, and exit flow distortions. A final geometry was chosen; this geometry will be built and tested by Rocketdyne. The results show that backward blade lean can be effective in red using the exit flow distortion of the impeller. Tip slots or holes were not beneficial because of the large inlet boundary layer. Also, it appears possible to raise the head coefficient of the baseline impeller without creating excessive flow distortion. The planned testing is necessary to verify the predictions of the flow code. / Master of Science

Thesis (MSc)--Stellenbosch University, 2014. / ENGLISH ABSTRACT: Grain regression analysis forms an essential part of solid rocket motor simulation. In this thesis a numerical grain regression analysis module is developed as an alternative to cumbersome and time consuming analytical methods. The surface regression is performed by the level-set method, a numerical interface advancement scheme. A novel approach to the integration of the surface area and volume of a numerical interface, as defined implicitly in a level-set framework, by means of Monte-Carlo integration is proposed. The grain regression module is directly coupled to a quasi -1D internal ballistics solver in an on-line fashion, in order to take into account the effects of spatially varying burn rate distributions. A multi-timescale approach is proposed for the direct coupling of the two solvers. / AFRIKAANSE OPSOMMING: Gryn regressie analise vorm ’n integrale deel van soliede vuurpylmotor simulasie. In hierdie tesis word ’n numeriese gryn regressie analise model, as ’n alternatief tot dikwels omslagtige en tydrowende analitiese metodes, ontwikkel. Die oppervlak regressie word deur die vlak-set metode, ’n numeriese koppelvlak beweging skema uitgevoer. ’n Nuwe benadering tot die integrasie van die buite-oppervlakte en volume van ’n implisiete numeriese koppelvlak in ’n vlakset raamwerk, deur middel van Monte Carlo-integrasie word voorgestel. Die gryn regressie model word direk en aanlyn aan ’n kwasi-1D interne ballistiek model gekoppel, ten einde die uitwerking van ruimtelik-wisselende brand-koers in ag te neem. ’n Multi-tydskaal benadering word voorgestel vir die direkte koppeling van die twee modelle.

As conventional chemical rockets reach the outer limits of their abilities, significant research is going into alternative thruster technologies, some of which decouple the maximum thrust and efficiency from the propellant's internal chemical energy by supplying energy to the propellant as needed. Of particular interest and potential is the electrically powered thruster, which promises very high specific thrust using relatively inexpensive and stable propellant gasses. Some such thrusters, specifically ion thrusters, have achieved significant popularity for various applications. However, there exist other classes of electrical thrusters which promise even higher levels of efficiency and performance. This thesis will focus on one such thruster type - the magnetoplasmadynamic thruster - which uses an ionized propellant flow and large currents to accelerate the propellant gas by electrical and magnetic force interactions. The necessary background will be presented in order to understand and characterize the operation of such devices, and a theoretical model will be developed in order to estimate the levels of performance which can be expected. Simulations will be performed and analyzed in order to better understand the principles on which these devices are designed. Finally, a thruster package will be designed and built in order to test the performance of the device and accuracy of the model. This will include a high-current power supply, ignition circuit, gas delivery system, and nozzle. Finally, the measured performance of this thruster package will be measured and compared to the theoretical predictions in order to validate the models constructed for this type of thruster.

High combustion temperatures and long operation durations require the use of cooling techniques in liquid propellant rocket engines. For high-pressure and high-thrust rocket engines, regenerative cooling is the most preferred cooling method. In regenerative cooling, a coolant flows through passages formed either by constructing the chamber liner from tubes or by milling channels in a solid liner. Traditionally, approximately square cross sectional channels have been used. However, recent studies have shown that by increasing the coolant channel height-to-width aspect ratio and changing the cross sectional area in non-critical regions for heat flux, the rocket combustion chamber gas side wall temperature can be reduced significantly without an increase in the coolant pressure drop. In this study, the regenerative cooling of a liquid propellant rocket engine has been numerically simulated. The engine has been modeled to operate on a LOX/Kerosene mixture at a chamber pressure of 60 bar with 300 kN thrust and kerosene is considered as the coolant. A numerical investigation was performed to determine the effect of different aspect ratio cooling channels and different number of cooling channels on gas-side wall and coolant temperature and pressure drop in cooling channel.

Thesis (MEng)--Stellenbosch University, 2014. / ENGLISH ABSTRACT: Recent years have seen a decline in the rate of space exploration due to the inefficiency of chemical rockets. Therefore alternative fuel efficient propulsion methods are being sought to enable cost effective deep space exploration. The high fuel efficiency of electric thrusters enable a spacecraft to travel further, faster and cheaper than any other propulsion technology available. Thus electric propulsion has become the propulsion of choice for scientists and engineers. A typically electric thruster contains some sort of electrode to ionise the propellant. Although this is feasible for short space missions, it becomes impractical for more ambitious space missions as electrodes erode over time. The alternative is to ionise the propellant using electromagnetic fields, which eliminates lifespan issues associated with electrode based thrusters. In order to examine methods of improving the lifespan and performance of electric thrusters, this thesis aimed to study the method of microwave discharge ionisation for an electric thruster. This includes the design of an RF Ion Thruster with extraction and acceleration grids to generate thrust. A 600 W 2.45 GHz magnetron (obtained from a conventional microwave oven), coupled to circular TM010 resonant cavity, was used to ionise neutral argon gas. The process of electron cyclotron resonance (ECR) was used to ensure the efficient ionisation of a high density plasma. The thrust was achieved with a three-grid system biased at high voltages to accelerate positively charged argon ions to high exhaust velocities. Results yielded the success of the designed electromagnetic based thruster, measuring approximatively 1.78 mN of thrust with a specific impulse of Isp = 3786 seconds. The ECR process produced a high plasma density with a plasma absorption rate of approximately 77% of the total input microwave power. The final results obtained were found to match the predicted results extremely well and resembled results found in literature. This demonstrates the efficiency of the RF ion thruster that was designed in this project and the future use in space exploration activities. However, future research needs to be undertaken on a controlled feedback system that will ensure optimal operating conditions for maximum performance. In addition, the method of grid-less acceleration needs to be studied to achieve maximum thrust and specific impulse. / AFRIKAANSE OPSOMMING: In onlangse jare het ’n afname in die tempo van die verkenning van die ruimte dit te danke aan die ondoeltreffendheid van chemiese vuurpyle. Derhalwe moet alternatiewe brandstof aandrywing metodes ondersoek word, om koste-effektiewe diep ruimte-eksplorasie moontlik te maak. Die hoë brandstof-doeltreffendheid van elektriese ontbranders stel ’n ruimtetuig in staat om verder, vinniger en goedkoper te reis as enige ander aandrywing tegnologie wat tans beskikbaar is. Dus het elektriese aandrywing metodes die aandrywings keuse vir wetenskaplikes en ingenieurs geword. ’n Tipies elektriese vuurpyl/aandrywer bevat ’n vorm van elektrode om die brandstof (argon gas) te ioniseer. Alhoewel hierdie elektrode proses van ionisasie effektief is vir kort ruimte missies, word dit onprakties vir meer ambisieuse ruimte missies as gevolg van verweering van elektrodes met verloop van tyd. ’n Alternatief is om die dryfmiddel/brandstof te ioniseer deur gebruik te maak van elektromagnetiese velde. Die elekromagnetiese velde sal die lewensduur van die vuurpyl vermeerder deur die verweering van elektrodes, wat geassosieer word met tipiese elektrieses vuurpyle, te elimineer. Hierdie tesis se doelwit is om die metode van mikrogolf ontslag ionisasie vir ’n elektriese vuurpyl/aandrywer te bestudeer om ten einde die lewensduur en doeltreffendheid van elektriese vuurpyl/aandrywer te ondersoek. Dit sluit in die ontwerp van ’n radio frekewensie ioon vuurpyl/aandrywer met ’n ontginning en versnelling matriks/rooster om stukrag te genereer. ’n 2,45 GHz magnetron (verkry vanaf ’n konvensionele mikrogolfoond), gekoppel aan ’n TM010 resonante holte, was gebruik om neutrale argon gas te ioniseer. Die proses van elektron siklotron resonansie (ESR) was gebruik om die doeltreffende ionisasie van ’n hoë digtheid plasma te verseker. Die aandrywing/stukrag was behaal met ’n drie-matriks-stelsel, bevoordeel deur hoë spannings om die positief-gelaaide argon ione te versnel. Resultate opgelewer, het die sukses van die ontwerp van ’n elektromagnetiese gebaseerde vuurpyl/aandrywer met ’n benaderde meting van ongeveer 1.78 mN van stukrag/aandrywing met ’n spesifieke impuls van Isp = 3786 sekondes bewys. Die ECR proses het ’n hoë plasma digtheid geproduseer met ’n plasma opname persentasie van ongeveer 77% van die totale inset mikrogolf energie. Die finale uitslae wat verkry was, het bevind dat die voorspelde resultate baie goed inpas met resultate in beskikbare literatuur. Dit dui op die doeltreffendheid van die RF ioon vuurpyl/aandrywer wat ontwerp is in hierdie projek vir die toekomstige gebruik in ruimte eksplorasie-aktiwiteite. Toekomstige navorsing moet op ’n beheerde terugvoer sisteem onderneem word, wat optimale werktoestande verseker vir maksimum prestasie. Daarbenewens moet die metode van matriks-lose versnelling bestudeer word, om maksimum versnelling/stukrag en spesifieke impuls te verseker.