Thesis (Professional Degree)--University of Missouri, School of Mines and Metallurgy, 1927. / The entire thesis text is included in file. Typescript. Title from title screen of thesis/dissertation PDF file (viewed October 5, 2009) Includes bibliographical references (p. 30-31) and index (p. 32-33).

Rocketry employs cryogenic refrigeration to increase the density of propellants, such as oxygen, and stores the propellant as a liquid. In addition to propellant liquefaction, cryogenic refrigeration can also conserve propellant and provide propellant subcooling and densification. Previous studies analyzed vapor conditioning of a cryogenic propellant, which occurred by either a heat exchanger positioned in the vapor or by using the vapor as the working fluid in a refrigeration cycle. This study analyzes the refrigeration effects of a heat exchanger located beneath the vapor-liquid interface of liquid oxygen. This study predicts the mass liquefaction rate and heat transfer coefficient for liquid oxygen using two different models, a Kinetic Theory Model and a Cold Plate Model, and compares both models to experimental data. The Kinetic Theory Model overestimated the liquefaction rate and heat transfer coefficient by five to six orders of magnitude, while the Cold Plate Model underestimated the liquefaction rate and heat transfer coefficient by one to two orders of magnitude. This study also suggested a model to predict the densification rate of liquid oxygen, while the system is maintained at constant pressure. The densification rate model is based on transient heat conduction analysis and provides reasonable results when compared to experimental data.

The purpose of this investigation was to determine the optimum conditions for explosion of a mixture of aluminum powder and liquid oxygen when fired without the use of a fixed detonator.

A magnetic fluid system could potentially replace mechanically moving parts in a satellite as a means of increasing system reliability and mission lifetime, but rather than a standard ferrofluid with magnetic particles, liquid oxygen (LOX) may be a more adequate working fluid. As a pure paramagnetic cryogen, LOX is already heavily used in space, but still requires basic research before being integrated into system development. The objectives of the research conducted were to verify LOX as a magnetic working fluid through experiment and establish a theoretical model to describe its behavior. This dissertation presents the theoretical, experimental, and numerical results of a slug of LOX being pulsed by a 1.1 T solenoid in a quartz tube with an inner diameter of 1.9 mm. The slug oscillated about the solenoid at 6-8 Hz, producing a pressure change of up to 1.2 kPa. System efficiency based on the Mason number was also studied for various geometric setups, and, using a one-dimensional, finite-differenced model in Matlab 2008a, the numerical analyses confirmed the theoretical model. The research provides groundwork for future applied studies with complex designs.

No description available.

L’oxygène liquide est utilisé principalement dans le domaine de la propulsion spatiale et la connaissance des mécanismes d’usure dans cet environnement est donc indispensable pour le développement des différents moteurs. Ce domaine est mal connu car l’oxygène liquide est un des rares fluides à associer trois propriétés spécifiques pouvant influencer les mécanismes du contact. Ces trois propriétés tribologiques spécifiques sont l’état liquide, la réactivité et la température cryogénique. Une campagne d’essais a été réalisée dans le cadre du projet européen In Space Propulsion-1 (ISP-1) afin d’identifier l’influence de chacune de ces propriétés sur un contact PCTFE/métal. Plusieurs comportements ont ainsi été explicités. D’abord, la présence de lubrification limite a été mise en évidence dans le cas d’un contact en azote liquide. Ensuite,la rugosité de la piste s’est révélé un paramètre fortement influent sur l’usure, les frottements ainsi que sur la formation d’un film de transfert de PCTFE sur le disque. Ce transfert de PCTFE a une grande influence sur le contact. Son épaisseur ainsi que sa régularité influencent directement les mécanismes du contact et plus particulièrement l’usure. Enfin, ces films de transfert se forment rapidement et leur épaisseur augmente avec la distance de glissement, changeant ainsi la vitesse d’usure. Tous ces mécanismes sont très dépendants de la température de surface au niveau du contact qui peut modifier les paramètres des matériaux. L’étude a donc été complétée en comparant une évaluation théorique de la température de surface avec une extrapolation de cette même température à partir des données mesurées dans le pion pendant la réalisation des essais. / Liquid oxygen is mainly used for space propulsion. The knowledge of the wear mechanisms in this environment is therefore essential for the development of the engines. Wear mechanisms in liquid oxygen are not well known because liquid oxygen is one of the few fluids combining three tribological properties that can influence contact mechanisms. These three specific tribological properties are the liquid state, the reactivity and the cryogenic temperature. A test campaign was performed in the frame of the European project In Space Propulsion-1 (ISP-1) to identify the influence of each one of these properties on the PCTFE/metal contact. Several behaviors have been explained. First, boundary lubrication has been demonstrated for contactin liquid nitrogen. Then, we confirmed that disk roughness was greatly affecting wear, friction and PCTFE transfer film formation on the disk. This PCTFE transfer film has a great influence on the contact properties. Its thickness and its regularity directly influence contact mechanisms, especially wear. Finally, the transfer film is easily formed and the thickness increases with the sliding distance, thereby changing the wear rate. All these mechanisms are highly dependent on the surface temperature at the contact interface, which can modify the materials parameters.The study was completed by comparing a theoretical evaluation of the surface temperature with an extrapolation of this same temperature from the measured data in the pin during the experiments.

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.

The present report aims to show the development process of a liquid level sensing system for the use within the propellant tanks of a launch vehicle. Whereas many recent launchers use discrete level sensing systems to indicate the amount of fuel on defined spots like fill-stop and engine cut-off, in this report a continuous capacitance based sensor probe is described. An accurate knowledge of the propellant level can serve as input for the throttling of the propellant valves to account for changes in the oxidizer-fuel ratio and therefore helps optimising the operation of the rocket engine. Furthermore, the more precise the amount of fuel is known, the less unused propellant mass will remain allowing a cost and mass optimisation for each mission. Additionally to the increased vehicle performance, the sensor designs described aim to have comparably low mass and cost. At the beginning of the report, an overview of different level measuring techniques is given before going into the special conditions and requirements regarding launch vehicles. Afterwards, the design and testing with RP-1 and LN2 of two different sensor probes using a capacitive measuring principle is described and compared to analytical calculations and numerical simulations using COMSOL Multiphysics. At the end, design suggestions for a flight probe and possible improvements for a higher reliability are given. All tested sensor designs show an accuracy of a few millimetres when tested within a settled, non-sloshing fluid. However, the theoretical models show a substantial deviation to the test data. / Föreliggande rapport syftar att visa utvecklingsprocessen av ett vätskenivåavkänningssystem för användningen i drivmedelstankarna i en bärraket. Medan många nya bärraketer använder diskreta nivåavkänningssytem för att ange mängden bränsle på definierade händelser som textit tankstopp och textit motor avstängning, beskrivs i denna rapport en kontinuerlig kapacitansbaserad sensor. En exakt kunskap om bränslenivån kan fungera som inmatning för strypningen av bränsleventilerna för att ta hänsyn till förändringar i förhållandet mellan oxidationsmedel och bränsle förhållandet. Detta hjälper därför till att optimera driften av raketmotorn. Dessutom, ju exaktare mängden bränsle är känd, desto mindre oanvänd bränsle återstår, vilket möjliggör en kostnads- och massoptimiering för varje uppdrag. Förutom den ökade fordonsprestandan syftar de beskrivna sensordesignerna till att ha jämförelsevis låg massa och kostnad. I början av rapporten ges en översikt över olika nivåmättekniker innan de speciella förhållandena och kraven för bärraketer berörs. Därefter beskrivs design och testning med med RP-1 och LN2 av två olika sensorer som använder en kapacitiv mätprincip och jämförs med analytiska beräkningar och numeriska simuleringar med COMSOL Multiphysics. I slutet ges designförslag för en flygprob och möjliga förbättringar för en högre tillförlitlighet. Alla testade sensorkonstruktioner visar en noggrannhet på några milimeter när de testas med en stilla, icke-plaskande vätska. De teoretiska modellerna visar dock en väsentlig avvikelse från testdatan.

A compact air separator demonstrator based on centrifugally enhanced distillation has been studied. The full size device is meant to be used on board of a Two Stage To Orbit vehicle launcher. The air separation system must be able to extract oxygen in highly concentrated liquid form (LEA, Liquid Enriched Air) from atmospheric air. The LEA is stored before being used in a subsequent rocket propulsion phase by the second stage of the launcher. Two reference vehicles are defined, one with a subsonic first stage and one with a supersonic first stage. In both cases, oxygen collection is performed during a cruise phase (M 0.7 and M 2.5 respectively). The aim of the project is to demonstrate the feasibility of the air separation system, investigate the separation cycle design, and assess that the separator design selected is suitable for the reference vehicles. The project is described from original base ideas to design, construction, extended testing and analysis of experimental results. Preliminary computations for a realistic layout have been performed and the motivations for the choices made during the process are explained. Test rig design, separator design and technical discussion are provided for a subscale pilot unit. Mass transport parameters and flooding limits have been estimated and experimentally measured. Performance has been assessed and shown to be sufficient for the reference Two Stage To Orbit vehicles. The technology developed is found suitable without further optimization, although some volume and mass reduction would be desirable for the supersonic first stage concept. There are many ways of optimisation that can be further investigated. The aim of this program, however, is not to fully optimize the device, but to demonstrate that a device based on a simple, robust, low-risk design is already suitable for the launch vehicles. On top of that analysis, directions for improvements are suggested and their potentials estimated. A complete assessment of those improvements requires further maturation of the technological concept through further testing and practical implementations. Directions for future work, general conclusions and a vehicle development roadmap have also been provided.

Vortex Tube (VT) is a simple device having no moving mechanical parts, in which compressed gas at high pressure is injected through one or more tangential nozzles into a vortex chamber resulting in the separation of the inlet flow into two low pressure streams. One of the streams is the peripheral flow that is warmer than the inlet stream while the other is the central (core) flow that is colder than the inlet stream. This separation of the inlet flow into high and low temperature streams is known as temperature or energy separation. It is suggested by many investigators that compressed air of few atmospheres pressure and at room temperature can produce temperatures as high as +200ºC at the hot end (peripheral flow exit) and as low as -50ºC at the cold end (core flow exit) of the VT. Though VTs have large potential for simple heating and cooling applications, the mechanism of energy separation is not clear so far. Based on their studies, many investigators have suggested various theories, different from each other, but having specific lacunas and is an unresolved issue. Also, till date, experimental and industrial designs of the VTs are based purely on empirical correlations. Apart from heating and cooling applications, VTs can also be used for separation of binary gas mixtures and separation of oxygen from two-phase precooled air stream. The conceptual futuristic cryogenic launch vehicle designs are being attempted with in-flight liquid oxygen (LOX) collection system that significantly improves the pay load fraction. Vortex tube technology is one of the few promising technologies for futuristic in-flight LOX separation based launch vehicles. This technology has significant advantages over its counterparts as it is a simple, compact and light weight, and most importantly have no moving parts and unaffected by gravity and orientation. In order that VTs become an acceptable technology for in-flight LOX separation system, it is necessary to achieve minimum oxygen purity of 90% with more than 60% yield (separation efficiency) for the oxygen enriched stream in the VT. A survey of the available open literature has shown very little reported details, in particular, on achieving the required specifications for in-flight LOX separation systems. Till date, the highest LOX purity of 60% with 40% separation efficiency has been reported with VT technology. In view of the above mentioned facts, the work carried out has been focused on to: • Optimize the critical parameters of the VT to achieve maximum energy separation by CFD and experimental studies. • Understand the flow behaviour in the VT by estimating the velocity, temperature and pressure profiles at various locations in the VT and validation of secondary circulation flow and its effect on the performance of energy separation in VT. • Estimation of the energy transfer between the core and the peripheral layers of fluid flow in VT by analytical and CFD methods to propose the most appropriate mechanism of energy separation in VT. • Design and development of a dedicated experimental setup for both energy separation and LOX separation studies in VTs. • Design and fabrication of straight and conical VTs and experimental programme on energy separation and LOX separation. • Development of the VT air separation technology to achieve the required specifications of in-flight LOX separation system for futuristic launch vehicles. With these specific objectives and motivations, the total work was carried out with the following planned and sequential steps: • The first step was the CFD modeling of the VT with the available CFD software (Star-CD) and obtain the energy separation phenomena for a 12mm diameter VT. After gaining sufficient confidence level, optimization of the critical parameters like the air injection nozzle profile, number of nozzles, cold end orifice diameter dc, length to diameter (L/D) ratio, hot gas fraction etc of the VT was carried out through CFD and experimental studies. • The studies show that 6 convergent nozzles perform better in comparison to other configurations like circular helical, rectangular helical, 2 convergent and 6 straight nozzles. The studies also show that cold end orifice diameter (dc) plays an important role on energy separation and bring out the existence of secondary circulation flow with improper design of cold end orifice diameter. Through our studies, the effect of cold end diameter on the secondary circulation flow has been evaluated for the first time. Also, the mechanism of energy transfer in VT based on heat pump mechanism enabled by secondary circulation flow as suggested by some investigators has been evaluated in our studies. The studies show that cold end orifice diameter dc = 7mm is optimum for 12mm diameter VT, which matches fairly with the correlations given by other investigators. The studies confirms that CFD modeling carried out in this work is capable of selecting the correct dc value for a VT, without resorting to the empirical correlations as a design guide or a laborious experimental programme. • Through the CFD and experimental studies on different length to diameter (L/D) ratios and hot gas fractions, maximum hot gas temperature of 391K was obtained for L/D = 30 with hot gas fraction of 12-15 % and minimum cold gas temperature of 267K for L/D = 35 was obtained for cold gas fraction ≈ 60% (lowest cold gas fraction possible with the present experimental system). • CFD analysis has been carried out to investigate the variation of static and total temperatures, static and total pressures as well as the velocity components of the particles as it progresses in the flow field, starting from the entry through the nozzles to the exit of the VT by tracking the particles to understand the flow phenomenon and energy transfer mechanism inside the VT. The studies indicate that the mechanism of energy transfer from the core flow to the peripheral flow in VT is predominantly occurs by the tangential shear work. Thus the investigations reported in the thesis have given a clear understanding of the contributing mechanism for energy separation in VT, which has been an unresolved issue for long time. The net energy transfer between the core and the peripheral fluid has been calculated analytically and compared with the values obtained by CFD model for VTs of L/D ratios equal to 10 and 30. The net energy transfer by analytical and CFD model for VT with L/D = 10 is 159.87W and 154.2W respectively whereas the net energy transfer by analytical and CFD model for VT with L/D = 30 is 199.87W and 192.3W respectively. The results show that CFD results are in very good agreement with the analytical results and CFD can be used as a tool for optimization of the critical parameters and to analyze the flow parameters and heat transfer analysis for VTs. Also, the net energy transfer between the core and peripheral fluids calculated analytically matches very well with that of the net energy transfer by CFD analysis, without considering the effect of acoustic streaming. Thus acoustic streaming may not be the mechanism of energy separation in VT as suggested by some investigators. • By optimizing the critical parameters of the 12mm diameter straight VT through CFD and experimental studies, LOX separation studies have been carried out using both straight and conical VTs of dc = 7mm and of different L/D ratios for high LOX purity and separation efficiency. It is observed that conical (3º divergence) VTs perform better as compared to straight VTs for LOX separation whereas straight VTs perform better for energy separation. The better performance of conical VT as compared to straight VTs can be attributed to its increased surface area for condensation-evaporation phenomenon of oxygen and nitrogen molecules. Experimental studies have been conducted to evaluate the influence of the inlet pressure and the inlet temperature (liquid fraction) on LOX purity. Studies indicate that for achieving high LOX purity for the studied experimental system, the inlet pressure is to be in the range of 6-6.5bar and there exists a very narrow band of inlet temperature zone in which high LOX purity can be achieved. Experimental studies on VTs show that VT can be optimized suitably either for high LOX purity with low separation efficiency or low LOX purity with high separation efficiency by adjusting the hot end mass fraction accordingly. It is also observed that it is not possible to obtain both high purity and high separation efficiency simultaneously with the single VT. Staging approach has to be adapted to achieve higher LOX purity with higher separation efficiency. By staging the VTs, the enriched air stream (hot end outlet flow) from the first stage of VTs is introduced to the inlet of the second stage of VTs. Experimental studies have been conducted to evaluate the design parameters on staging of VTs. LOX purity of 48% with 89% separation efficiency has been achieved for conical first stage VT of L/D = 25. LOX purity of about 94% with separation efficiency of 84% has been achieved for 50% oxygen content at the inlet of the second stage VT. Similarly, LOX purity of 96% with separation efficiency of 73.5% has been achieved for 60% oxygen content at the inlet of the VT. This is the highest LOX purity and separation efficiency reported so far indicating that, conical VT of optimized diameter, L/D ratio and orifice diameter can yield the hot end flow very close to the target value of futuristic in-flight LOX separation based launch vehicles. The present investigation has focused the optimization of the critical parameters of VTs through CFD and experimental studies. It has also given an insight to the mechanism of energy transfer between the core and peripheral flow in VT by evaluating two of the existing theories on mechanism of energy transfer in VT. The studies also highlighted the fact that custom designed and precision fabricated VTs can be very useful for obtaining maximum / minimum temperatures of fluid flow as well as LOX separation with high purity and high separation efficiency needed for futuristic in-flight LOX separation based space launch vehicles.