No description available.

No description available.

A jet ejector may be used as a compressor or to enhance thrust of watercraft or aircraft. Optimization of jet ejectors as compressors and thrust augmenters was conducted using the software GAMBIT (Computer Aided Engineering (CAE) tool for geometry and mesh generation) and FLUENT (Computational Fluid Dynamics (CFD) solver kit). Scripting languages PYTHON and SCHEME were used to automate this process. The CFD model employed 2D axis symmetric, steady-state flow using the ε−k method (including wall functions) to model turbulence. Initially, non-dimensionalization of the jet ejector as a gas compressor was performed with respect to scale, fluid, and operating pressure. Surprisingly, rather than the conventional parameters like Mach or Re number, the results showed a completely new parameter (christenedGM- Gauge Mach) that when kept constant will result in non-dimensionalization. Non-dimensionalization of a jet ejector for watercraft propulsion was conducted using 2D axis symmetric, steady-state flow modeling using the ε−kmethod (including wall functions). It showed consistent results for the same velocity ratio (r) of nozzle velocity to free-stream velocity for different scales, fluids, and ambient pressures.

A jet ejector may be used as a compressor or to enhance thrust of watercraft or aircraft. Optimization of jet ejectors as compressors and thrust augmenters was conducted using the software GAMBIT (Computer Aided Engineering (CAE) tool for geometry and mesh generation) and FLUENT (Computational Fluid Dynamics (CFD) solver kit). Scripting languages PYTHON and SCHEME were used to automate this process. The CFD model employed 2D axis symmetric, steady-state flow using the ε−k method (including wall functions) to model turbulence. Initially, non-dimensionalization of the jet ejector as a gas compressor was performed with respect to scale, fluid, and operating pressure. Surprisingly, rather than the conventional parameters like Mach or Re number, the results showed a completely new parameter (christenedGM- Gauge Mach) that when kept constant will result in non-dimensionalization. Non-dimensionalization of a jet ejector for watercraft propulsion was conducted using 2D axis symmetric, steady-state flow modeling using the ε−kmethod (including wall functions). It showed consistent results for the same velocity ratio (r) of nozzle velocity to free-stream velocity for different scales, fluids, and ambient pressures.

A supersonic ejector is used to pump a secondary gas using a supersonic primary gas flow by augmentation of momentum and energy in a variable area duct. The internal compressible flow through an ejector has many complex gas dynamic features, like compressible shear layers and associated shock interactions. In many practical applications, ejectors are operated in the choked flow regimes where higher operating pressure ratios and mass flow rates are encountered. On the other hand, rather low entrainment and subsonic secondary flow dynamics (referred as the mixed regime of operation) dominate the dilution and purging applications of ejectors. The fundamental understanding of the flow dynamics associated with gaseous mixing process in the ejector especially in the mixed operational regime is still unclear. Obtaining a comprehensive understanding of the flow through a supersonic ejector in the mixed regime through experimental investigations is the prime focus of the present study. A new supersonic ejector test facility is designed, fabricated and established in the laboratory during the course of this study. The effects of using different gases in the secondary flow have been investigated. Two novel methods to improve the ejector by enhancing mixing are also implemented. Optical diagnostic tools (Time-resolved Schlieren and laser scattering) and wall static pressure measurements are used to investigate the dynamics of mixing process inside the ejector. State of the art image processing codes are developed to determine the length in the ejector for which the primary and the secondary flows are separate, referred here as the non-mixed length from the results of the flow visualization studies. Exhaustive experiments are carried out on the two dimensional rectangular supersonic ejector by varying the mass flow rates of primary and secondary flows, primary stagnation pressure, for two locations of the nozzle in the ejector. The non-mixed length determined from quantitative flow visualization tools is found to lie within 4.5 to 5.2 times the height of the duct (20 mm). The non-mixed flow length determined from flow visualization studies corroborates well with the wall static pressure measurements. A significant reduction of non-mixed length of about 46.7% is caused by shock wave-boundary layer interactions in the supersonic nozzle at over-expanded conditions. Further, the effects of differences in molecular weight and ratio of specific heats on the performance are also studied using cylindrical supersonic ejector at low entrainment ratios (0.008 to 0.06). In these studies air is used as the primary fluid while argon and helium are used in the secondary flow segment of the ejector. The results indicate that Argon has better entrainment characteristics compared to helium. Two novel supersonic nozzles (the tip rig nozzle and Elliptic Sharp Tipped Shallow lobed nozzle) are also devel- oped to enhance mixing in the ejector. About 30% enhancement of entrainment ratio is observed with the newly designed nozzle geometries. Illustrative numerical simulations are also carried out to complement the experimental studies.

No description available.

Cette thèse vise à étudier la valorisation des rejets thermiques de basse température (inférieure à 100 °C) par les machines à absorption, dont principalement les transformateurs de chaleur à absorption (AHT) pour générer de la chaleur à une température plus élevée, et les cycles de réfrigération par absorption (ARC) pour la production de frigories. Les performances des machines à absorption sont exprimées suivant les températures et les COP qu’elles peuvent atteindre. Ces deux paramètres dépendent de l’architecture de la machine, des composants utilisés et de la paire de fluides réfrigérant-absorbant circulant dans la machine. L’objectif principal de ce travail est une nouvelle génération d’AHT qui puisse rendre utile la chaleur avec une différence de température par rapport à la source de chaleur disponible (rejet thermique) d’au moins 50°C. Pour l‘ARC, on cible une nouvelle génération de cycles capables de produire des frigories à une température inférieure à -20°C.Les résultats obtenus représentent une avancée majeure dans le domaine des cycles à absorption. En effet, les objectifs fixés de valorisation de chaleur basse température, jusqu’à 45°C, en rendant disponible de la chaleur haute température (supérieure à 120°C) et basse température (inférieure à -20°C) ont été atteints. Ceci représente une différence de température entre le rejet thermique et la chaleur utile de 75°C minimum. L’utilisation d’éjecteurs avec certains fluides de travail a permis le développement d'une nouvelle génération de cycles à absorption. / This thesis aims at studying the low temperature waste heat recovery (less than 100°C) by the use of absorption machines, mainly absorption heat transformers (AHT) to generate heat at a higher temperature, and absorption refrigeration cycles (ARC) to generate chilling power. The performances of absorption machines are expressed according to the temperatures and the COPs which they can reach. These two parameters depend on the cycle configuration, the components used and the refrigerant-absorbent pair of fluids circulating in the machine. The main objective of this work is a new generation of AHT which can produce heat with a difference in temperature compared to the available heat source (heat rejection) of at least 50°C. For the ARC, a new generation of cycles capable of producing chilling power at temperatures lower than -20°C is aimed.The results obtained represent a major progress in the absorption cycles field. The objectives set for this work of low temperature heat recovery, up to 45°C, by making available high temperature heat (above 120°C) and low temperature (below -20°C) have been achieved. This represents a minimal temperature difference between the waste heat and the useful heat of 75°C. The use of ejectors with specific working fluids allowed the development of a new generation of absorption cycles.

<p>  </p> <p>High altitude test facilities are needed for ground testing of upper stage rocket engines or small satellite thrusters with high expansion ratio nozzles to ensure full-flowing nozzle conditions. Rocket exhaust diffusers and ejector systems are essential components of these facilities and are frequently used to set desired simulated altitude/low pressure conditions and pump out rocket exhaust products. </p> <p>This dissertation combined experimental and numerical efforts on diffuser-ejector systems. The experimental efforts included the development of a Second Throat Exhaust Diffuser (STED) to aid with the qualification of space thrusters in the Purdue Altitude Chamber Facility. While performing these experiments, we characterized the single and two-stage ejector systems operating in conjunction with the diffuser to obtain and maintain specific simulated altitudes. </p> <p>The concurrent numerical effort focused on validating a Computational Fluid Dynamics (CFD) approach based on Reynolds-averaged Navier–Stokes equations flow simulations. After validating the ejector CFD, we used it to derive a corrective coefficient of a lumped parameter ejector model (LPM) developed previously for the ejectors used in the Purdue Altitude Facility. We created variable coefficient maps for the stages of the two-stage ejector system using the same LPM and the test data from one of our experiments. </p> <p>We designed, manufactured, and then validated a STED for altitude testing of a ~50 lbf hypergolic hybrid motor as a part of a NASA JPL project. The designed STED enabled the operation of the hybrid motor for the full duration of the test firing (about 2 seconds) at a simulated altitude of 102,000 feet, slightly above the targeted altitude of 100,000 feet. We also validated our diffuser CFD approach by creating a simulation using the measured diffuser back pressure and the average motor chamber pressure. </p> <p>We then devised an experiment to investigate several diffuser–ejector system configurations using cold gas thrusters with conical and bell nozzles. The main aim of that experiment was to explore the effects of different thruster nozzle geometries, diffuser geometries, and thruster/ejector operational parameters on the performance of a diffuser–ejector system. For all the configurations tested, we reported on the minimum starting and operating pressure ratios and corresponding correction factors on the normal shock method. The large hysteresis regions obtained mostly with bell nozzles having a high initial expansion angle provided an opportunity to economize the facility resources. In some cases which were later found to violate STED second throat contraction limits, we experienced a choking flow at the second throat. Then, we studied the second throat contraction limits in detail using CFD in addition to the experimental data and explored minimum diffuser second throats enabling diffuser starting and improving aerodynamic efficiency. </p> <p>Finally, we machined a larger scale cold gas thruster with different nozzle geometries (having throat diameters in the range of 0.367 – 0.52 inches) from acrylic rods to study possible flow separation and gas condensation events that could occur during tests in the altitude chamber. The main difference here with the previous experiment was that the diffuser (JPL STED) was fixed, and the two-stage ejector system was used to create the necessary back pressure. With the experiments performed at varying axial gaps between the nozzle exit and diffuser inlet, we were able to investigate the effect of that on the diffuser performance. The experimental data collected in this work and the complementary numerical efforts served to generate the operating envelope of the Purdue Altitude Chamber Facility.  </p>

Rapid mixing and spreading of supersonic jets are two important characteristics in supersonic ejectors, noise reduction in jets and fuel mixing in supersonic combustion. It helps in changing the acoustic and thermal signature in supersonic exhaust. The supersonic nozzles in most cases result in compressible mixing layers. The subsonic nozzles form incompressible mixing layers but at high Mach numbers even they form compressible mixing layers. Compressible mixing layers have been found to have much lower mixing and spreading rates than incompressible mixing layer Birch & Eggers (1972). In order to enhance the spreading and mixing of mixing layers from supersonic nozzles various active and passive methods have been deviced. Active methods include fluid injection, fluid lobes and plasma actuation. Passive methods are mostly based on modifying the nozzle geometry such that the fluid expansion is ideal or the shock cell is broken. Many nozzles with exotic shapes have been developed to obtain mixing enhancements in supersonic jets Gutmark et al. (1995). To achieve enhanced mixing an innovative nozzle named as the Elliptic Sharp Tipped Shallow (ESTS) lobed nozzle has been developed in L.H.S.R., I.I.Sc., India Rao & Jagadeesh (2014). This nozzle has a unique geometry involving elliptical lobes and sharp tips. These lobes are generated using a simple manufacturing process from the throat to the exit. This lobed and sharp tipped structure introduces stream wise vortices and azimuthal velocity components which must help in enhanced mixing and spreading. The ESTS lobed nozzle has shown mixing enhancement with 4 lobes. The spreading rate was found to be double of the reference conical nozzle. This thesis is motivated by the need to investigate the flow physics involved in the ESTS lobed nozzle. The effect of varying the number of lobes and the design Mach number of the nozzle on the mixing and spreading characteristics will be further discussed. Visualisation studies have been performed. The schlieren and planar LASER Mie scattering techniques have been used to probe the flow. Instantaneous images were taken at axial planes with the reference conical and ESTS nozzles with three, four, five and six lobes. The nozzles are for design Mach number 2.0 and 2.5. The stagnation chamber pressure was maintained to obtain over expanded, ideally expanded and under expanded flows. LASER scattering was obtained by seeding the flow with water to observe the behaviour of the primary flow. The condensation of moisture due to the cold primary flow mixing with the ambient air was exploited to scatter laser and observe the flow structures in the mixing layer. A comparison of the images of the reference conical nozzle and the ESTS lobed nozzles shows changes in the mixing layers due to the ESTS lobed nozzles. The image of the reference conical nozzle shows a distinct potential core and mixing layers all along the length of the image. For the ESTS lobed nozzles this distinction becomes unclear shortly after the nozzle exit. Thus mixing of the primary flow and ambient air is seen to be enhanced in the case of all the ESTS lobed nozzles. The flow in the case of the ESTS lobed nozzles if found to be highly non axis symmetric. The starting process of the nozzles has been visualised using time resolved schlieren. Image processing was performed on the nozzles to quantify the spread rate. The shock structure of the nozzles has been studied and found to be modified due to the lobed geometry. The level of convolution of the mixing layer due to the lobed structure has been studied using fractal analysis. The four lobed nozzle was found to have the highest spread rate and th most convoluted shear layer. Hence this nozzle was further studied using background oriented schlieren and particle image velocimetry to quantify the flow field. These experimental results have been compared with CFD simulations using the commercial software CFX5. The computations and experiments don’t match accurately but the trends match. This allows for simulations to be used as a good first approximation. The acoustic properties of a jet are dependent on the flow structure behaviour. The ESTS lobes have been found to change the flow structure. Hence the ESTS lobed nozzle was predicted to change the acoustic signature of the flow. The acoustic measurements of the flow were carried out at National Aerospace Laboratories, Bengaluru. The screech of the overexpanded flow was seen to be eliminated and the overall sound levels were found to have been reduced in all cases. Thus the lobed nozzle was found to have acoustic benefits over the reference conical nozzle. Thus the ESTS lobed nozzle has been studied and compared with the conical nozzle using several methods. The changes due to the lobed structure have been studied quantitatively. Future studies would focus on the change in thrust due to the lobed structure. Also new geometries have been proposed inspired by the current design but with possible thrust benefits or manufacturing benefits.