Hybrid Aircraft Propulsion System Synthesis (HAPSS) is a computer program that sizes and analyzes pure-series hybrid electric propulsion systems for aircraft. The development of this program began during a NASA SBIR contract, in conjunction with Empirical Systems Aerospace (ESAero), with the creation of a propulsion fan design tool. Since the completion of this contract in July 2010, the HAPSS program has been expanded to combine the many aspects of a hybrid propulsion system such as the propulsive fans, electric motors, generators, and controllers, and the internal combustion engines. This thesis describes the benefits and drawbacks of aircraft hybrid propulsion systems to reveal the usefulness of a program of this nature. The methodology behind HAPSS, the creation of the program, its operation, and its many applications are also discussed in detail. Finally, this thesis includes a brief example in which HAPSS is used to analyze a hybrid propulsion system for a commercial transport aircraft. This example demonstrated the usefulness of the program and revealed interesting behavior and trends unique to hybrid propulsion. To date, the HAPSS program has been utilized on several different contract projects in which an aircraft hybrid propulsion system was designed. In the summer of 2012, a government organization in conjunction with ESAero will begin funding a contract to continue the development of HAPSS by adding functionality and improving accuracy while making the tool available to other government agencies.

Experiments were conducted on the Cal Poly air augmented rocket (AAR) in order to characterize two-dimensional flowfield phenomenon occurring in the mixing duct. The testing utilized a direct connect system where high pressure nitrogen is fed into the combustion chamber, to form a primary flow. The high pressure nitrogen is then expanded through a nozzle, with an area ratio of 22 and an exit area of 0.75 in2, up to Mach 4.3. Secondary air is entrained from a plenum chamber which is used to create a lower stagnation pressure for the secondary flow. The two flows mix in a duct that has a cross sectional area of 2.06 in2. The maximum pressure ratio, the ratio of primary to secondary stagnation pressure, achieved during testing was 132. The stagnation pressures of the primary and secondary flows are transient throughout the test. The quasi-steady portion of each run increased with increasing pressure ratio. Pressure and temperature measurements were collected from ten test runs. Shadowgraph images were taken of the mixing duct during testing in order to image the interactions between the primary and secondary flows. The images show an oblique shock forming in the primary flow. The angle of the shock matches theoretical predictions to within 8.41%. The oblique shock begins at a distance of 1.5 inches downstream of nozzle exit when the AAR is operating in the Fabri choked condition. The images also show the mixing region which forms between the primary and secondary flows. The mixing region represents as much as 25% of the cross-sectional area of the flow field in the mixing duct two inches downstream of the nozzle exit. An analysis of the secondary Mach number in the mixing duct shows that Fabri choking is occurring during testing. The secondary Mach number decreases as pressure ratio increases, in the Fabri choked condition. The transition to Fabri choking occurs at a pressure ratio of 100, suggesting that this is the pressure ratio of the saturated case. The shape of the primary plume was compared to results from a 2-D simulation developed to predict the flow field inside the Cal Poly AAR. Although, the simulation is unable to predict the entire flowfield, modifications made it able to predict the velocity of the secondary, entrained, flow within 3.7%. The modified simulation also predicts the that the primary plume will have expanded 98% of its total distance from the centerline of the mixing duct 1.7 inches downstream of the primary nozzle exit. Pressure data taken along the wall of the mixing duct was used to identify the location of Fabri choking in the mixing duct. Tests showed that Fabri choking is occurring between 1 inch and 2.5 inches downstream of the nozzle exit. The location of Fabri choking moves farther downstream of the nozzle as pressure ratio increases.

No description available.

The need for micropropulsion solutions for spacecraft has been steadily increasing as scientific payloads require higher accuracy maneuvers and as the use of small form-factor spacecraft such as CubeSats becomes more common. Of the technologies used for this purpose, electrospray thrusters offer performance that make them an ideal choice. Electrosprays offer high accuracy impulse bits at low power and high efficiency, and have low volume requirements. Design choice reasoning and preliminary testing results are presented for two electrospray thruster designs. The first thruster, named the Demonstration thruster, is operated in atmospheric conditions and serves as a highly visible example of the basic concepts of electrospray technology applied to micropropulsion. It features a single capillary needle emitter and the acetone propellant flow is driven actively by a syringe pump. The second thruster, named the Research thruster, is operated in the vacuum environment and is designed for modularity for its expected use in future research efforts. Propellant flow is also driven actively using a syringe pump. Initial configuration of the Research thruster is a linear array of five capillary needle emitters, though testing is conducted with only one emitter in this thesis. Tests using un-doped glycerol and sodium iodide doped glycerol (20% by weight) are conducted for the Research thruster. Both thruster designs use stainless steel 18 gauge blunt dispensing needles (0.038 in / 0.965 mm ID) as their emitters. Applied voltage to the emitter(s) relative to the grounded extractor is swept from 2100 V to 3700 V for the Demonstration thruster testing and from 4000 V to 4500 V for the Research thruster. Currents incident on a collection plate downstream of the emission plume and on the extractors of the thrusters were measured directly with a pico-ammeter. Measurements made during testing of the Demonstration thruster are inconsistent due to charge loss as propellant travels through the air, though currents as high as 5.1x10-9 A on the collection plate and 2x10-7 A on the extractor are recorded. Currents for Research thruster testing using un-doped glycerol were measured as high as 4.9x10-8 A on the collection plate and 5x10-9 A on the extractor, showing an interception rate as high as 17%. Currents using sodium iodide doped glycerol were measured as high as 7x10-7 A on the collection plate. Discussion is given for the visual qualities of cone-jet emission for all testing. Keywords:

With an increased demand in Cube Satellite (CubeSat) development for low cost science and exploration missions, a push for the development of micro-propulsion technology has emerged, which seeks to increase CubeSat capabilities for novel mission concepts. One type of micro-propulsion system currently under development, known as Pocket Rocket, is an electrothermal plasma micro-thruster. Pocket Rocket uses a capacitively coupled plasma, generated by radio-frequency, in order to provide neutral gas heating via ion-neutral collisions within a gas discharge tube. When compared to a cold-gas thruster of similar size, this gas heating mechanism allows Pocket Rocket to increase the exit thermal velocity of its gaseous propellant for increased thrust. Previous experimental work has only investigated use of the gas discharge tube's orifice for propellant expansion into vacuum. This thesis aims to answer if Pocket Rocket may see an increase in thrust with the addition of a micro-nozzle, placed at the end of the gas discharge tube. With the addition of a conical ε = 10, α = 30° micro-nozzle, performance increases of up to 6% during plasma operation, and 25% during cold gas operation, have been observed. Propellant heating has also been observed to increase by up to 60 K within the gas discharge tube.

Les cerf-volants existent depuis l'Antiquité, mais leur utilisation comme moyen de récupération de l'énergie éolienne est relativement récente. Pourtant, leur légèreté et leur capacité à aller chercher les vents forts et réguliers en altitude en font un dispositif compétitif pour produire de l'électricité ou pour tracter des navires commerciaux. En effet, un cerf-volant peut espérer produire plus de $10$ kW.m$^{-2}$. Cela implique qu'un cerf-volant de $1000$ m$^2$ pourrait apporter une assistance substantielle (typiquement $20$ $%$) à la propulsion des plus gros cargos actuels. Cette thèse s'intéresse à deux problèmes associés au développement de tels cerf-volants:Comment les faire décoller et atterrir de manière autonome et sans risque de les perdre? L'utilisation de cerf-volants à structure gonflable donne l'avantage d'avoir une aile rigide et légère en vol et compacte lors de son stockage. Pour aider au dimensionnement de ces cerf-volants, nous étudions dans le première partie de la thèse le comportement des structures gonflables soumis à des chargements statiques et dynamiques.Comment s'assurer de son vol stable? Une fois qu'il a décollé, un cerf-volant doit pouvoir rester en l'air. Il apparait cependant que dans certaines conditions, les cerf-volants entrent dans des oscillations de grandes amplitudes avant de tomber au sol. Grâce à des expériences en soufflerie, nous étudions dans la seconde partie de la thèse l'origine de ces oscillations et les conditions à réunir pour les éviter. / Kites exist since ancient times, but their use as wind energy harvesting device is relatively recent. Still, their light weight and ability catch strong and steady winds in altitude make them a competitive mean to generate electricity or to tow commercial ships. Indeed, a kite can typically produce $10$ kW.m$^{-2}$. This implies that a $1000$ m$^2$ kite could provide substantial assistance ($20$ $%$) to the propulsion of the biggest current tankers. This thesis focuses on two issues associated with the development of such kites:How can one perform autonomous take off and landing without the risk of losing them? Kites with inflatable structures take advantage rigidity and lightness during flight and from high compactness during storage. It also allows them to float if they crash on the ocean. To design those kites, we study in the first part of the thesis the behavior of inflatable structures under static and dynamic loadings.How can one achieve a stable flight? Once it takes off, it appears that under certain conditions, the kites undergo large amplitude oscillations that eventually lead to their fall onto the ground. Using wind tunnel experiments, we examine in the second part of the thesis the origin of these oscillations and the conditions which prevent them from occurring.

Thesis (M.Eng.), Memorial University of Newfoundland, 2000. / Bibliography: leaves 101-105.

This work focuses on the concept of sustainable propellant collection. The concept consists of gathering ambient gas while on-orbit and using it as propellant. Propellant collection could potentially enable operation in very-low Earth orbits without compromising spacecraft lifetime. This work conducts a detailed analysis of propellant collection from a physics perspective in order to test the assertions of previous researchers that propellant collection can dramatically reduce the cost of propellant on-orbit. Major design factors for propellant collection are identified from the fundamental propellant collection equations, which are derived in this work from first principles. A sensitivity analysis on the parameters in these equations determines the relative importance of each parameter to the overall performance of a propellant-collecting vehicle. The propellant collection equations enable the study of where propellant collection is technically feasible as a function of orbit and vehicle performance parameters. Two case studies conducted for a very-low Earth orbit science mission and a propellant depot-type mission serve to demonstrate the application of the propellant collection equations derived in this work. The results of this work show where propellant collection is technically feasible for a wide range of orbit and vehicle performance parameters. Propellant collection can support very-low Earth operation with presently available technology, and a number of research developments can further extend propellant-collecting concepts' ability to operate at low altitudes. However, propellant collection is not presently suitable for propellant depot applications due to limitations in power.

Thesis (M.Eng.)--Memorial University of Newfoundland, 2004. / Bibliography: leaves 177-180.

La caractérisation et la simulation des écoulements internes au sein des moteurs àpropergol solide, en considérant des mécanismes physiques fortement couplés, constituentl’objectif principal de ce mémoire de thèse. Dans cette optique, la conjonction entrefluide/régression de surface/couplage fluide structure a imposé de déployer une stratégiepropre lors du développement de la modélisation numérique. En effet, le modèle intègre untraitement de frontière immergée couplé avec un suivi de frontière mobile afin de pouvoirrendre compte de la formidable variation géométrique interne subie au cours d’un tir. Côtéfluide, un maillage automatique est nécessaire et la gestion de ce dernier s’appuie sur undéveloppement récursif avec structure hiérarchisée de type 2n tree. Une attention particulièrea été portée sur le solveur lui-même avec une approche explicite en temps et un schémanumérique basé sur l’approche de Roe avec limiteur de flux au second ordre. Des cas testsont été réalisés afin de valider le solveur et les différentes conditions aux limites introduites,notamment des conditions spécifiques développées pour les besoins de simulation. Lespremiers résultats soulignent tout l’intérêt du modèle proposé et sauf erreur de notre part,pour la première fois, l’analyse des sources tourbillonnaires responsables des instabilités ausein de ces moteurs a été étudiée en intégrant les effets du changement continu de géométrie.Finalement, la faisabilité d’une interaction forte entre solveur fluide et solveur solide a étéréalisée sur un modèle simplifié d’un moteur segmenté.L’ensemble des développements permet un accès aux mécanismes couplés complexeset aux fortes interactions au sein des moteurs à propergol solide et offre de nouvellesperspectives dans la caractérisation des mécanismes fortement couplés. / Characterization and simulation of internaI flow within the solid rocket motors, considering the physicalmechanisms strongly coupled, are the main focus of this thesis objective. In this context, the conjunctionbetween fluid/regression surface/fluid coupling structure imposed deploy c1ean during the development ofnumerical modeling strategy. Indeed, the model incorporates treatment coupled with an immersed boundarytracking moving boundary in order to realize the tremendous internai geometric variation experienced during ashot. Fluid side, an automatic mesh is required and the management of the latter is based on a recursivehierarchical structure development with type 2" tree. Particular attention was paid to the solver itself with anexplicit approach to time and a numerical scheme based on the approach of Roe with flow limiter in the secondorder. Tests cases were conducted to validate the sol ver and different boundary conditions introduced, inc1udingspecific conditions developed for the purpose of simulation. The first results emphasize the interest of theproposed and unless our error model, for the first time, the analysis of the sources responsible vortex instabilitiesin these engines has been studied by incorporating the effects of continuous change in geometry. Finally, thefeasibility of a strong interaction between fluid and solid solver was conducted on a simplified model of a multiengine.AlI the developments allows access to complex mechanisms coupled and strong interactions in solidrocket motors and off ers new insights into the characterization of strongly coupled mechanisms.