The Task Scheduling Paradigm is a general technique for leveraging fine and coarse grain parallelism from applications of several domains with minimum impact on code readability, relying on the automatic inference of data dependencies among tasks. The performance of Task Parallel applications is correlated with the speed at which the underlying Task Scheduling System is able to detect such dependencies, something that is critical for fine-granularity workloads, which cannot amortize scheduling overheads with long periods of useful computation. That being the case, several groups have recently been developing FPGA-accelerated Task Scheduling Systems architectures where a software Task Scheduling Runtime is able to offload its bookkeeping computations to an FPGA-based accelerator with the goal of efficiently scheduling fine-grained tasks to CPU cores. Even though these FPGA-accelerated systems offer substantial gains over the software-only baseline, it is also true that FPGA-CPU communication bottlenecks prevent such designs from handling scenarios with either large number of cores or very fine-grained tasks. With that in mind, we proposed the implementation of a Native Task Scheduling System that is, a processor with native support for task scheduling embedded into its architecture with the goal of substantially reducing these overheads. More specifically, this project aimed at embedding the HW logic of Picos, a mature Task Scheduling Accelerator developed by the Barcelona Supercomputing Center (BSC), into Rocket Chip, an open-source, silicon-proven, multi-core implementation of RISC-V. The ISA of the resulting system provides special instructions for Task Applications to interact with this Task Scheduling Logic, ruling out all FPGA-CPU communication latencies. To evaluate the prototype performance, we both (1) adapted Nanos, a mature Task Scheduling runtime, to benefit from the new task-scheduling-accelerating instructions; and (2) developed Phentos, a new HW-accelerated light weight Task Scheduling runtime. Our experiments show that task parallel programs using Nanos-RV the Nanos version ported to our system are on average 2.13 times faster than those being serviced by baseline Nanos, while programs running on Phentos are 13.19 times faster, considering geometric means. Using eight cores, Nanos-RV is able to deliver speedups with respect to serial execution of up to 5.62 times, while Phentos produces speedups of up to 5.72 times. / Paralelismo por Tarefas é uma técnica genérica de extração de paralelismo de granularidade arbitrária aplicável a programas de vários domínios, com mínimo impacto sobre legibilidade de código, baseada na inferência automática de dependências de dados entre tarefas. O desempenho de aplicações paralelas baseadas nesse paradigma depende da velocidade com a qual o runtime de Paralelismo por Tarefas que lhe dá suporte é capaz de detectar tais dependências, fato que é ainda mais crítico para aplicações envolvendo tarefas de granularidade fina, já que nesse cenário o overhead de escalonamento não é amortizado por períodos significativamente maiores de computação útil. Recentemente, diversos grupos têm desenvolvido Sistemas de Suporte a Paralelismo por Tarefas acelerados por FPGAs, os quais são capazes de fazer offload das operações de inferência de dependências para um acelerador em FPGA de modo a melhorar o seu desempenho ao lidar com tarefas de granularidade fina. Por outro lado, ainda que esses sistemas acelerados por FPGA apresentem ganhos substanciais com relação às alternativas baseadas puramente em software, o desempenho dessas soluções é prejudicado por gargalos de comunicação entre a CPU e a FPGA, os quais limitam a capacidade desses sistemas de lidar com cenários envolvendo grande número de núcleos ou tarefas muito finas. Motivados por isso, implementamos um Sistema de Suporte Nativo a Paralelismo por Tarefas isto é, um processador com suporte arquitetural nativo a Paralelismo por Tarefas com o objetivo de reduzir consideravelmente tais overheads de comunicação. Mais especificamente, integramos a lógica em hardware do Picos, um acelerador de Paralelismo por Tarefas desenvolvido pelo Barcelona Supercomputing Center (BSC), ao Rocket Chip, uma implementação multi-core de código livre do RISC-V desenvolvida pela Universidade da Califórnia, Berkeley. O sistema resultante contém em sua ISA (Instruction Set Architecture) as instruções necessárias para que aplicações baseadas em tarefas possam interagir diretamente com essa lógica de escalonamento, minimizando os overheads associados ao uso de runtimes intermediários e eliminando toda a latência de comunicação FPGA-CPU. Para avaliar a performance do protótipo que então se construiu, nós tanto (1) adaptamos o runtime de escalonamento de tarefas Nanos para que ele pudesse ser acelerado pelas novas instruções de escalonamento de tarefas, quanto (2) criamos um novo runtime leve de escalonamento de tarefas a que demos o nome de Phentos. Nossos experimentos mostram que programas baseados em paralelismo por tarefas usando o runtime Nanos-RV a versão do runtime Nanos com suporte ao sistema que produzimos são executados em média 2,13 vezes mais rapidamente do que versões dos mesmos programas utilizando a versão básica do Nanos, enquanto programas executados com o Phentos são em média 13,19 vezes mais rápidos do que suas versões correspondentes baseadas na mesma versão básica do Nanos. Tais valores médios correspondem à média geométrica dos conjuntos de dados pertinentes. Usando oito núcleos, Nanos-RV entrega ganhos de desempenho com relação a execuções seriais de até 5,62 vezes, enquanto Phentos entrega ganhos de até 5,72 vezes.

Rocket-Based Combined Cycle (RBCC) engines combine the best performance characteristics of air-breathing systems such as ramjets and scramjets with rockets with the goal of increasing payload/structure and propellant performance and thus making LEO more readily accessible. The idea of using RBCC engines for Single-Stage-To-Orbit (SSTO) trans-atmospheric acceleration is not new, but has been known for decades. Unfortunately, the availability of detailed models of RBCC engines is scarce. This thesis addresses the issue through the construction of an analytical performance model of an ejector rocket in a dual combustion propulsion system (ERIDANUS) RBCC engine. This performance model along with an atmospheric model, created using MATLAB was designed to be a preliminary `proof-of-concept' which provides details on the performance and behavior of an RBCC engine in the context of use during trans-atmospheric acceleration, and also to investigate the possibility of improving propellant performance above that of conventional rocket powered systems. ERIDANUS behaves as a thrust augmented rocket in low speed flight, as a ramjet in supersonic flight, a scramjet in hypersonic flight, and as a pure rocket near orbital speeds and altitudes. A simulation of the ERIDANUS RBCC engine's flight through the atmosphere in the presence of changing atmospheric conditions was performed. The performance code solves one-dimensional compressible flow equations while using the stream thrust control volume method at each station component (e.g. diffuser, burner, and nozzle) in all modes of operation to analyze the performance of the ERIDANUS RBCC engine. Plots of the performance metrics of interest including specific impulse, specific thrust, thrust specific fuel consumption, and overall efficiency were produced. These plots are used as a gage to measure the behavior of the ERIDANUS propulsion system as it accelerates towards LEO. A mission averaged specific impulse of 1080 seconds was calculated from the ERIDANUS code, reducing the required propellant mass to 65% of the gross lift off weight (GLOW), thus increasing the mass available for the payload and structure to 35% of the GLOW. Validation of the ERIDANUS RBCC concept was performed by comparing it with other known RBCC propulsion models. Good correlation exists between the ERIDANUS model and the other models. This indicates that the ERIDANUS RBCC is a viable candidate propulsion system for a one-stage trans-atmospheric accelerator.

Les lanceurs ont un impact sur la composition de l'atmosphere, et en particulier sur l'ozone stratospherique. Parmi tous les types de propulsion, les moteurs à propergol solide ont fait l'objet d'une attention particulière car leurs émissions sont responsables d'un appauvrissement significatif d'ozone dans le panache des lanceurs lors des premières heures suivant le lancement. Ce phénomène est principalement dû à la conversion de l'acide chlorhydrique, un composé chimique présent en grandes quantités dans les émissions de ce type de moteur, en chlore actif qui réagit par la suite avec l'ozone dans un cycle catalytique similaire à celui responsable du "trou de la couche d'ozone", cette diminution périodique de l'ozone en Antarctique. Cette conversion se produit dans le panache supersonique, où les hautes températures favorisent une seconde combustion entre certaines espèces chimiques du panache et l'air ambiant. L'objectif de cette étude est d'évaluer la concentration de chlore actif dans le panache d'un moteur à propergol solide en utilisant la technique des Simulations aux Grandes Echelles (SGE). Le gaz est injecté à travers la tuyère d'un moteur et une méthode de couplage entre deux instances du solveur de mécanique des fluides est utilisée pour étendre autant que possible le domaine de calcul derrière la tuyère (jusqu'à l'équivalent de 400 diamètres de sortie de la tuyère). Cette méthodologie est validée par une première SGE sans chimie, en analysant les caractéristiques de l'écoulement supersonique avec co-écoulement obtenu par ce calcul. Ensuite, le chimie mettant en jeu la conversion des espèces chlorées a été étudiée au moyen d'un modèle "hors-ligne" permettant de résoudre une chimie complexe le long de lignes de courant extraites d'un écoulement moyenné dans le temps résultant du calcul précédent (non réactif). Enfin, une SGE multi-espèces est réalisée, incluant un schéma chimique auparavant réduit afin de limiter le coût de calcul. Cette simulation représente une des toutes premières SGE d'un jet supersonique réactif, incluant la tuyère, effectuée sur un domaine de calcul aussi long. En capturant avec précision le mélange du panache avec l'air ambiant ainsi que les interactions entre turbulence et combustion, la technique des simulations aux grandes échelles offre une évaluation des concentrations des espèces chimiques dans le jet d'une precision inédite. Ces résultats peuvent être utilisés pour initialiser des calculs atmosphériques sur de plus larges domaines, afin de modéliser les réactions entre chlore actif et ozone et de quantifier l'appauvrissement en ozone dans le panache. / Rockets have an impact on the chemical composition of the atmosphere, and particularly on stratospheric ozone. Among all types of propulsion, Solid-Rocket Motors (SRMs) have given rise to concerns since their emissions are responsible for a severe decrease in ozone concentration in the rocket plume during the first hours after a launch. The main source of ozone depletion is due to the conversion of hydrogen chloride, a chemical compound emitted in large quantities by ammonium perchlorate based propellants, into active chlorine compounds, which then react with ozone in a destructive catalytic cycle, similar to those responsible for the Antartic "Ozone hole". This conversion occurs in the hot, supersonic exhaust plume, as part of a strong second combustion between chemical species of the plume and air. The objective of this study is to evaluate the active chlorine concentration in the far-field plume of a solid-rocket motor using large-eddy simulations (LES). The gas is injected through the entire nozzle of the SRM and a local time-stepping method based on coupling multi-instances of the fluid solver is used to extend the computational domain up to 400 nozzle exit diameters downstream of the nozzle exit. The methodology is validated for a non-reactive case by analyzing the flow characteristics of the resulting supersonic co-flowing under-expanded jet. Then the chemistry of chlorine is studied off-line using a complex chemistry solver applied on trajectories extracted from the LES time-averaged flow-field. Finally, the online chemistry is analyzed by means of the multi-species version of the LES solver using a reduced chemical scheme. To the best of our knowledge, this represents one of the first LES of a reactive supersonic jet, including nozzle geometry, performed over such a long computational domain. By capturing the effect of mixing of the exhaust plume with ambient air and the interactions between turbulence and combustion, LES offers an evaluation of chemical species distribution in the SRM plume with an unprecedented accuracy. These results can be used to initialize atmospheric simulations on larger domains, in order to model the chemical reactions between active chlorine and ozone and to quantify the ozone loss in SRM plumes.

The design of a liquid rocket propulsion system, unlike that of a standalone system, is intertwined with the overall development of a number of associated systems and is influenced by a multitude of conditions and considerations: from the requirements needed to accomplish the mission to the rationalizations involved behind the development of each rocket system and/or component. In my thesis, the preliminary design of a “new generation” 30 kN rocket engine driven by an electric pump feed system and running on liquid methane and liquid oxygen is performed. The propulsion system would be employed on a hypothetical small-lift orbital-class twin-stage rocket to deliver a light payload of about 200 kg into a circular 500 km LEO. Such topics as the selection of bipropellant combinations, the feasibility of electric pump feed systems, design methodologies for thrust chambers, for nozzles in particular, management of the high thermal energy and the selection of compatible wall materials, as well as the design of an injector have been looked comprehensively into. It is realized that methalox is indeed better than both hydrolox (with regard to density impulse) and kerolox (in terms of specific impulse). Besides, a suite of attractive characteristics makes the bipropellant a combination of choice to power rockets of the future. Yet more notably, an electric-pump-fed engine cycle is, under the right circumstances of engine operation, established to outperform both the pressure feed system and the turbopump feed system. With constant advancement in battery technologies, improvement of both power density and energy density to achieve much higher performance is but a matter of time. The adoption of a propulsion system such as ours for a mission objective as outlined above, therefore, is not just viable but unquestionably realistic. Two thrust chamber versions—a sea-level variant for the booster stage and a vacuum-optimized variant for the upper stage—are developed for our rocket. And both the nozzles employ a TOP “thrust optimised parabolic” contour; also, the booster stage comprises a cluster of 9 engines in a parallel burn arrangement. Concerning thermal management, the entirety of the booster-stage thrust chamber implements regenerative cooling (using Inconel 625), whereas the aft of the upper-stage nozzle section implements radiative cooling (with Niobium C-103). Further, the injector faceplate (also of Inconel 625) comprises two concentric patterns of unlike impingement doublet sets: with 80 pairs on the outer ring and 40 pairs on the inner ring. With rational assumptions, our hypothetical launch vehicle is deemed to have a mass of roughly 17200 kg (200 kg of which is the payload) and a delta-v of approximately 9600 m/s—quite within the desirable range of specifications for small-lift orbital-class twin-stage rockets of today.

No description available.

The noise of a rocket engine is a complex and complicated phenomenon which has been studied for more than half a century [1]. There are many sources to this noise, but due to its great potential impact to the surrounding structures, including the vehicle itself, it is important to have a model to estimate the acoustic environment the engine produces. This estimation can be used for the design of a launch pad or a flame deflector. Such a model was developed and then tested, by measuring the noise levels with six microphones at Rocket Factory Augsburg´s Vertical Test Stand at Esrange Space Centre. Three out of six microphones yielded valuable data. A comparison between the estimated and measured noise was then conducted which showed similar trends. The peak frequency in the estimation was in the order of 1 kHz. A sensitivity study was made to investigate the difference in Sound Power Level (SPL) when the engine and test stand parameters were adjusted. The parameters with the greatest effect on the SPL are the Mach number, thrust, potential core length, and impingement distance. The difference in SPL between the estimation and measured noise is 0-20 dB with a lower difference at lower frequencies and a higher difference at higher frequencies. The difference was higher when comparing the estimation to the test with an overpowered engine, with differences of up to 20 dB higher than the estimation in the upper frequencies. Differences with nominal engine data was up to 15 dB higher than the estimation, constrained to lower frequencies. Above 30% of the peak frequency, the noise was consistently lower than the estimation. The estimation can be concluded to likely be conservative at higher frequencies, further testing or a new estimation is necessary with accurate engine data. / Ljudet från en raketmotor är ett komplext och komplicerat fenomen som har studerats i mer än ett halvt sekel [1].Det finns många källor till det ljudet, men på grund av risken att det skadar omgivande strukturer, inklusive raketen, är det viktigt att ha en modell för att estimera ljudmiljön motorn produceras. Estimeringen kan användas för att designa en uppskjutningsramp eller en flammdeflektor. En sådan modell var utvecklad och testad genom att göra ljudmätningar med sex mikrofoner på Rocket Factory Augsburg´s testanläggning på Esrange Space Centre. Tre av sex mikrofoner gav värdefull data. En jämförelse gjordes mellan det estimerade och uppmätta ljudet vilket visade liknande trender. Toppfrekvensen i estimeringen var i storleksordningen 1 kHz. En känslighetsstudie gjordes för att undersöka skillnaden i ljudnivån (SPL) när motorns och testanläggningens parametrar justerades. Parametrarna med störst påverkan på ljudnivån var Mach numret, drivkraft, längden av flamman och avståndet till deflektorn. Ljudskillnaden mellan det estimerade och uppmätta ljudet var mellan 0-20 dB med mindre skillnad på lägre frekvenser och större skillnad på högre frekvenser. Skillnaden var större vid jämförelse mellan estimering och testet med en kraftigare motor, med skillnader på upp till 20 dB över estimeringen på de högre frekvenserna. Skillnaderna för nominell motordata var upp till 15 dB högre än estimeringen, begränsat till lägre frekvenser. Över 30% av toppferkvensen var ljudet konsekvent lägre än estimeringen. Estimeringen kan sannolikt konstateras vara konservativ på högre frekvenser, ytterligare tester eller estimeringar behövs med exakt motordata.

Thesis (M.S.)--Mississippi State University. Department of Physics and Astronomy. / Title from title screen. Includes bibliographical references.

A visualisation procedure was developed which predicts excited hydroxyl (OH*) radiation from the Computational Fluid Dynamics (CFD) solutions of cryogenic hydrogen-oxygen rocket flames. The model of backward ray tracing through inhomogeneous media with a continuously changing refractive index was implemented. It obtains the optical paths of light rays that originate in the rocket chamber, pass through the window and enter a simulated camera. Through the use of spectral modelling, the emission and absorption spectra eλ and κλ are simulated on the ray path from information about temperature, pressure and concentration of constituent species at relevant points. By solving a radiative transfer equation with the integration of emission and absorption spectra along the ray line-by-line, a spectral radiance is calculated, multiplied with the spectral filter transmittance and then integrated into total radiance. The values of total radiances at the window edge are visualised as a simulated 2D image. Such images are comparable with the OH* measurement images. The modelling of refraction effects results in up to 20 % of total radiance range absolute difference compared to line-of-sight integration. The implementation of accurate self-absorption corrects significant over-prediction, which occurs if the flame is assumed to be optically thin. Modelling of refraction results in images with recognisable areas where the effect of a liquid oxygen (LOx) jet core can be observed, as the light is significantly refracted. The algorithm is parallelised and thus ready for use on big computational clusters. It uses partial pre-computation of spectra to reduce computational effort.

Physical mechanism for high frequency instability in cryogenic engines at low hydrogen injection temperature has been a subject of debate for long time. Experimental and early developmental studies revealed no instabilities and it was only much later when liquid hydrogen at lower initial temperature (~50 to 100 K) was injected into the combustion chamber that instabilities were detected. From the compilations of the experimental data related to the instability of cryogenic engines by Hulka and Hutt, it was found that the instability was strongly connected to the temperature of hydrogen. Experiments conducted with hydrogen temperature ramping from a higher value to lower values indicated that the temperatures in excess of 90 K favor stability under most practical operating conditions. Even though this has been known for over forty years, there has been no clear and simple explanation for this. Many physical mechanisms have been hypothesized to explain how temperature ramping causes instability, but all appear to have limited range of applicability. Current understanding of cryogenic engine combustion instability has been achieved through a combination of experimental investigation and approximate analytical models as well as CFD tools. Various researchers have tried to link the low hydrogen injection temperature combustion instability phenomena with various potential mechanisms for combustion instability. They involve coupling of combustion acoustics with atomization, vaporization, mixing, chemical kinetics or any combination of these processes. Various studies related to the effect of recess, injector hydrodynamics, acoustic damping of gas liquid scheme injectors and effect of drop size distribution on the stability characteristics of cryogenic engines were compiled in the thesis. Several researchers examined fuel droplet vaporization as the rate controlling mechanism. Recently a new method for the evaluation of stability characteristics of the engine using model chamber were proposed by Russians and this is based on mixing as the rate controlling mechanism. Pros and cons of this method were discussed. Some people examined the combustion instability of rocket engines based on chemistry dynamics. A considerable amount of analytical and numerical studies were carried out by various researchers for finding out the cause of combustion instability. Because of the limitations of their analysis, they could not successfully explain the cause of combustion instability at low hydrogen injection temperature. A compilation of previous numerical studies were carried out. A number of researchers have applied CFD in the study of combustion instabilities in liquid propellant rocket engines. In the present thesis, a theoretical model has been developed based on the vaporization of droplets to predict the stability characteristics of the engine. The proposed concept focuses on three dimensional simulation of combustion instability for giving some meaningful explanations for the experimental work presented in the literature. In the present study the pressure wave corresponding to the transverse modes were superimposed on a three dimensional steady state operating conditions. Steady state parameters were obtained from the three dimensional combustion modeling. The conservation equations for mass, momentum and energy are non dimensionalized for facilitating the order of magnitude analysis. In order to do the stability analysis, variables are represented as the sum of their steady values and deviation from the steady state. A harmonic time dependence is assumed for the perturbations. For the transverse mode of oscillations independent variables of the zeroth order equations are r and θ only and the dependant variables are not functions of the axial distance. The axial dependence comes only through the first order equations. In this analysis, the wave motion in the combustion chamber is assumed to be linear, confining the nonlinearity to the vaporization process only. The reason behind making this assumption is that the vaporization process is the major mechanism driving the instability. Vaporization histories of liquid oxygen drops in a combustor with superimposed transverse oscillations were computed and stability characteristics of the engine were estimated. The stability characteristics of the engine are accessed from the solutions of first order equations. Effects of various parameters like droplet diameter, hydrogen injection temperature and hydrogen injection area on the stability characteristics of cryogenic engines are studied. A comparison of predicted and published experimental results was made which showed general agreement between experiment and computation. The present study and experimental results show clearly that hydrogen injection velocity is the critical parameter for instability rather than hydrogen injection temperature. What has happened in actual experiments when hydrogen injection temperature is varied is an effective alteration of the injection velocity that leads to the situation of instability. For higher relative velocity between hydrogen and liquid oxygen, the response of the vaporization rate in the presence of pressure wave is minimum compared to lower relative velocity. Due to this cryogenic engines will go to unstable mode at lower relative velocity.

The main objective of this thesis is to introduce the reader to the problematics of air-launch and to a custom design solution applying this concept. The specifics of this method of bringing a payload into orbit are described and explained. Overview of projects which use aircraft to launch spacecraft is included. Determination of primary parameters of a launch vehicle designed to carry a payload of a specified mass is conducted. The required flight performance has been estimated, a computational model has been developed in software MATLAB, and a multidisciplinary optimization of the design parameters has been performed using a genetic algorithm optimization method. Parameters of the designed air-launched rocket are compared with those of a ground-launched rocket. According to the specific criteria, the Airbus A310-300 aircraft was selected as the most suitable transport aircraft to be used for launching the designed launch vehicle. The last part of the thesis is devoted to the proposal of necessary modifications and estimation of the flight performance.