A condição ótima de funcionamento de uma tubeira em um motor foguete com escoamento isentrópico, implica que a velocidade na garganta (seção de menor área) seja equivalente à velocidade do som local, condição de Mach 1 e bocal engasgado. Pode-se alcançar essa condição reduzindo a área da seção do escoamento até a área crítica, velocidade sônica. Após a garganta acontece a expansão e se alcança velocidades supersônicas no divergente. Para manter a condição de Mach 1 na garganta em motores foguetes, trabalha-se com pressões superiores à necessária para se engasgar o bocal. Isto ocorre porque tenta-se compensar instabilidades ou variações de volumes produzidos na combustão ou queima. Usando uma pressão de trabalho maior, impõe-se que a condição de Mach 1 fique mantida durante toda a queima do combustível, isso implica em usar tubos mais resistentes à pressão e maior massa do tubo-motor. Observou-se experimentalmente que em algumas situações construtivas se podem modificar a pressão e temperatura necessárias para engasgar o bocal aumentando o comprimento da garganta. O comprimento do estrangulamento pode estabelecer uma condição para formação e evolução da camada limite e esta condição restringir a área nominal, modificando o regime do escoamento. Um equipamento especialmente desenvolvido para esse ensaio compara resultados de cinco modelos de motores, divididos em dois grupos, cada grupo com áreas de entrada, garganta e saída iguais, porém com comprimentos diferentes de garganta. Em análise experimental, observou-se que a pressão de trabalho e a temperatura são influenciadas pelo comprimento da garganta, interferindo na relação entre as pressões internas e de garganta e apresentando condições de engasgamento mensuráveis. Essas medidas foram conduzidas no presente estudo de doutorado. / The optimum operational condition of a rocket motor nozzle with isentropic flow implies that the velocity at the throat (the section with smallest area) is equivalent to the speed of the local sound. This speed is also called Mach 1 and it is said that at this condition the nozzle is choking. One can achieve this condition by reducing the cross-sectional area of the flow to the critical area resulting in a sonic speed. Beyond the nozzle throat, in the divergent section of the motor, flow expansion occurs and reaches supersonic speeds. To maintain the condition of Mach 1 at the throat, higher pressures than the one necessary to choke the nozzle are applied. This practice is done in order to compensate for jitter or variations of volumes produced in the combustion process. Using a higher operating pressure guarantees that a Mach 1 speed is maintained throughout the combustion process. Consequently, due to this higher operating pressure, more resistant tubes are needed to withstand this higher pressure and an increase in the motor weight is inevitable. It was observed experimentally that some constructional modifications of the motor can alter the pressure and temperature required for choking. This was noted with increasing the bottleneck length of the nozzle throat which was able to establish a condition for the formation and evolution of the boundary layer, restricting the nominal area and thus modifying the flow regime. In this study, the results of five engine models are compared using a specially designed equipment. The rockets were divided into two groups, each with equal inlet, throat, and exit areas, but having different throat lengths. In experimental analysis, it was observed that the working pressure and temperature are influenced by the length of the throat, interfering in the relationship between the internal pressures and throat presenting measurable choking conditions which were conducted in this doctorate thesis study.

The manufacturing process for solventless extruded double base propellants involves a number of rolling and reworking stages. Throughout these processes a decrease in weight average molecular weight was observed, this was attributed to denitration. Differential scanning calorimetery data indicated that the reworking stages of extruded double base propellant manufacture were crucial to the homogenisation of the propellant mixture. To determine the homogeneity of the final extruded product, a sample was analysed across its diameter. No variations in stabiliser concentration, molecular weight, or Vickers hardness were detected. An accelerated thermal ageing trial simulating up to 8 years of ageing at 25°C was carried out to evaluate the storage characteristics. Reductions in stabiliser concentration, number average molecular weight, weight average molecular weight and polydispersity compared with un-aged samples were observed. The glass transition temperature measured using differential scanning calorimetery decreased by ~3°C. The decrease was attributed to the initial denitration reducing the energy of bond rotation and shortening the polymer chains, both factors reducing the energy required for movement. Modulus values determined from dynamic mechanical analysis temperature scanning experiments, did not detect significant variation between un-aged and aged samples. Though it was considered that variations would be likely if a more extensive ageing program was completed. In order to evaluate propellant behaviour at very high and low frequencies, time temperature superposition (TTS) and creep testing were carried out. The TTS technique superpositioned data well, allowing future investigation of high frequency propellant properties. Creep testing was considered to be an appropriate approach, though the equipment available was not optimised for such testing. This thesis is concerned with understanding how propellants are manufactured from nitrocellulose, nitroglycerine and other constituents. It is also about how the propellants decompose during long periods of time in storage, and how these changes can be measured using thermal and mechanical methods. It is about how the physical, chemical and thermal properties of the propellant composition change throughout the manufacture. This is relevant as it could be used to develop more efficient manufacturing processes, allow operators to adjust processes to tailor product properties or be used to re-design manufacturing to compensate for a different starting material. The thesis also considers how and why the properties of the product change over the course of years of storage. A specific focus on whether changes in mechanical and thermal properties occur, and if so how they can be detected. / © Cranfield University

Norwegian Centre for Space-related Education (NAROM) performs education within subject areas related to space, such as space physics, atmosphere and space technology. This thesis describes the development of a transmitter to be used in NAROM ́s studentrocket. The requirements were that the transmitter should send with FM at 2279,5 MHz with an output of at least 750 milliwatt and with a speed of 512 kbit/s when NRZ coding is used.The transmitter that was designed and tested in this project used an IC that modulated the signal to FM and sent it out on a low frequency. Then the frequency was multiplied by a IC to the correct value, 2279,5 MHz. Experiments showed that the transmitter worked satisfying, but a problem occurred: The frequency multiplier sent out unwanted overtones and the output power was to low. A final theoretical design was made to solve these problems. The final design uses the same modulation components as the tested transmitter, but filters are added and the amplifiers are changed to be able to handle the filters and give out a higher power. The transmitter can be use for other application that need a transmitter with bit-rate of up to 600 kbit/s and frequency range of 2250-2300 MHz without change of filters. If the filters is change the frequency range will be 2250-2550 MHz. / <p>Validerat; 20131025 (global_studentproject_submitter)</p>

Nanosatellites, such as Cubesats, are a rapidly growing sector of the space industry. Their popularity stems from their low development cost, short development cycle, and the widespread availability of COTS subsystems. Budget-conscious spacecraft designers are working to expand the range of missions that can be accomplished with nanosatellites, and a key area of development fueling this expansion is the creation of micropropulsion systems. One such system, originally developed at the Australian National University (ANU), is an electrothermal plasma thruster known as Pocket Rocket (PR). This device heats neutral propellant gas by exposing it to a Capacitively Coupled Plasma (CCP), then expels the heated gas to produce thrust. Significant work has gone towards understanding how PR creates and sustains a plasma and how this plasma heats the neutral gas. However, no research has been published on varying in the device's geometry. This thesis aims to observe how the size of the RF electrode affects PR operation, and to determine if it can be adjusted to improve performance. To this end, a thruster has been built which allows the geometry of the RF electrode to be easily varied. Measurements of the plasma density at the exit of this thruster with different sizes of electrode were then used to validate a Computational Fluid Dynamics (CFD) model capable of approximately reproducing experimental measurements from both this study and from the ANU team. From this CFD, the number of argon ions in the thruster was found for each geometry, since collisions between argon ions and neutrals are primarily responsible for the heating observed in the thruster. A geometry using a 10.5 mm electrode was observed to produce a 23% increase in the quantity of ions produced compared to the baseline 5 mm electrode size, and a 3.5 mm electrode appears to produce 88% more ions.

Data Week UPC 2021 día 1 / Data WeeK UPC es un evento anual organizado por las Facultades de Negocios e Ingeniería, con el propósito de reunir a investigadores y expertos en la gestión empresarial para reflexionar acerca del papel de la Ciencia de Datos en la generación de valor en las organizaciones. Nueve expositores de distintas instituciones se unirán a las 4 fechas del Data Week 2021 este 23, 25, 26 y 27 de noviembre, para reflexionar acerca de los retos en el proceso de la transformación de datos para la toma de decisiones. No se pierdan la oportunidad de participar en este espacio en el que discutiremos las principales tendencias en cuanto a la aplicación de la ciencia de datos en la gestión empresarial. 7:00 PM BUSINESS ANALYTICS: DEL ROCKET SCIENCE A UNA SOLUCIÓN DE NEGOCIOS El análisis correcto de los datos y el entendimiento de sus patrones tienen el potencial de aportar en la competitividad de una organización. En esta charla se analizarán algunas de las estrategias que permiten orientar los negocios hacia una cultura Data Driven. 8:00 PM TRANSFORMACIÓN DIGITAL La definición de una estrategia digital y el adecuado análisis y explotación de datos corporativos representan grandes retos para las organizaciones, en esta charla se abordará esta problemática y su relación con la transformación digital.

Hypergolic hybrid motors are advantageous for rocket propulsion due to their simplicity, reliability, low weight, and safety. Many hypergolic hybrid fuels with promising theoretical performance are not practical due to their sensitivity to temperature or moisture. Ammonia borane (AB) has been proposed and studied as a potential hypergolic hybrid fuel that provides both excellent performance and storability. This study investigates the effect of droplet impact velocity, pellet composition, and storage humidity on ignition delay of AB with white fuming nitric acid as the oxidizer. Most ignition delays measured were under 50 ms with many under 10 ms and some even under 2 ms, which is extremely short for hybrid systems. Higher droplet velocities led to slightly shorter ignition delays, and exposing samples to humidity slightly increased ignition delay. An AB pellet composition of at least 20% epoxy binder was found to minimize ignition delay. The epoxy facilitates ignition by absorbing or adhering the oxidizer and slowing the reaction with the fuel, preventing oxidizer expulsion and holding it close to the fuel. These results emphasize the importance of binder properties in hypergolic hybrids. Pellets varying in composition and storage method were extinguished and reignited with the oxidizer to demonstrate reignition capability.

Hybrid Rockets are popular in the aerospace industry due to their storage safety, simplicity, and controllability during rocket motor burn. However, they produce fuel regression rates typically 25% lower than solid fuel motors of the same thrust level. These lowered regression rates produce unacceptably high oxidizer-to-fuel (O/F) ratios that produce a potential for motor instability, nozzle erosion, and reduced motor duty cycles. To achieve O/F ratios that produce acceptable combustion charactersitics, traditional cylindrical fuel ports are fabricated with very long length-to-diameter ratios to increase the total burning area. these high aspect ratios produce further reduced fuel regression rate and trust levels, poor volumetric efficiency, and a potential for lateral structural loading issues during high thrust burns. In place of traditional cylindrical fuel ports, it is proposed that by researching the effects of centrifugal flow patterns introduced by embedded helical fuel port structures, a significant increase in fuel regression rates can be observed. The benefits of increasing volumetric efficiencies by lengthening the internal flow path will also be observed. The mechanisms of this increased fuel regression rate are driven by enhancing surface skin friction and reducing the effect of boundary layer "blowing" to enhance convective heat transfer to the fuel surface. Preliminary results using additive manufacturing to fabricate hybrid rocket fuel grains from acrylonitrile-butadiene-styrene (ABS) with embedded helical fuel port structures have been obtained, with burn-rate amplifications up to 3.0x than that of cylindrical fuel ports.

This research investigates the application of additive manufacturing techniques for fabricating hybrid rocket fuel grains composed of porous Acrylonitrile-butadiene-styrene impregnated with paraffin wax. The digitally manufactured ABS substrate provides mechanical support for the paraffin fuel material and serves as an additional fuel component. The embedded paraffin provides an enhanced fuel regression rate while having no detrimental effect on the thermodynamic burn properties of the fuel grain. Multiple fuel grains with various ABS-to-Paraffin mass ratios were fabricated and burned with nitrous oxide. Analytical predictions for end-to-end motor performance and fuel regression are compared against static test results. Baseline fuel grain regression calculations use an enthalpy balance energy analysis with the material and thermodynamic properties based on the mean paraffin/ABS mass fractions within the fuel grain. In support of these analytical comparisons, a novel method for propagating the fuel port burn surface was developed. In this modeling approach the fuel cross section grid is modeled as an image with white pixels representing the fuel and black pixels representing empty or burned grid cells. (80 pages)

Due to renewed interest in the sub orbital rocket program of the Florida Space Authority and a surplus of Super Loki Sounding Rockets, an effort to improve the usefulness of this surplus is herein undertaken. Currently, the capacity of the payload compartment in the upper stage of the Super Loki system is very limited. A redesign of the upper stage will allow larger and more versatile payloads to be carried, assuming the appropriate design parameters are met. It has therefore been undertaken to create a design procedure that is comprehensive in scope in order to affect this redesign. This procedure includes five major components. These are the separation of the upper and lower stages, the stability of the vehicle, the altitude and velocity of the rocket, the mechanical loading and finally the aerodynamic heating. Semi-empirical methods were used whenever possible to allow comparison with experimental data. This procedure revealed that larger diameter upper stages might be used up to a reasonable maximum of four inches. The four-inch modification is found to be stable as were the smaller modifications considered. The altitude and velocity of the rocket were found via a simple Eulerian time stepping scheme resulting in an estimate of approximately 148,000ft for the four-inch dart. The mechanical loading analysis allowed for the material selection for the rocket components. Reinforced steel fins and carbon fiber tubing, for the payload section, are adequate to meet expected mechanical loads, those being, 16000psi for the fin section due to launcher forces, 22800psi for compressive plus torsion forces on the composite section and 18000psi for the ejection stresses. An ablative coating is considered necessary to counteract the 760ºF temperatures along the composite tube.

The design, development, integration and testing of a full-featured, Lab-Scale Hybrid Rocket Engine was not only envisioned to be the chosen method of putting student payloads into space, but to be an invaluable teaching resource. The subject of the present thesis is the analysis, design, development, integration and demonstration of a lab-scale hybrid rocket motor. The overarching goal of this project was to establish a working developmental lab model from which further research can be accomplished. The lab model was specifically designed to use a fuel source that could be studied in normal laboratory conditions. As such, the rocket engine was designed to use Hydroxyl Terminated Polybutadiene as the fuel and Liquid Nitrous Oxide as the oxidizer. Developing the rocket engine required the usage of several electronics modules and a software package. The custom-designed electronics modules were a Signal Conditioning & Data Amplification Interface and a Data Acquisition Network. The software package was coded in Visual Basic (VB). A MathCAD regression rate computer model was designed and written to geometrically constrain the engine design. Further, the computer model allowed for the "what-if" situations to be evaluated. Using ProPep, solutions to the Equilibrium Thermodynamics Equations for the fuel and oxidizer mixture were obtained. The resultants were used as initial input to the computer model for predicting the lab-scale rocket's Chamber Pressure, Chamber Temperature, Ratio of Specific Heats and Molecular Weight. Details on the model, the rocket hardware, and the successful test firing are provided.