Dans les domaines de la réhabilitation, des sciences du sport et de l'ergonomie, l'évaluation des capacités de génération de force (CGF) peut aider à mieux comprendre les capacités motrices humaines. Le but de cette thèse a été d'évaluer les CGF du membre supérieur prédites au moyen de deux types de formalismes. Le premier provient du domaine de la robotique et a été utilisé pour déterminer l'ellipsoïde de force normalisé (EFN) et biomécanique (EFB), le polytope de force normalisé (PFN) et biomécanique (PFB). Pour une posture, ils sont calculés à partir d’un modèle polyarticulé du membre supérieur et de données sur les couples articulaires isométriques maximaux. Le second type fait appel à un modèle musculosquelettique afin de modéliser les CGF sous la forme d’un polytope de forces (PFMS). Tous ces modèles ont été comparés à un polytope de forces mesurées (PFM). Pour le construire, les forces maximales isométriques exercées par le membre supérieur au niveau de la main ont été évaluées dans vingt-six directions différentes. Enfin, le PFMS a été appliqué dans le cadre de la propulsion en fauteuil roulant afin de caractériser l'application des forces lors de cette tâche et un nouvel indice d’évaluation de la performance postural (IPP) a été proposé. / In fields like rehabilitation, sports sciences and ergonomics, the evaluation of the force feasible set (FFS) of the human limbs may help to better understand the human motor abilities. The aim of this thesis was to compare the upper-limb force capacity at the hand predicted by two different kinds of FFS formalism. The first one originating from the robotics field was used to compute the force ellipsoid (FE), scaled force ellipsoid (SFE), force polytope (FP) and scaled force polytope (SFP). For one posture, they are computed from the upper-limb model and hypotheses and data on maximum isometric joint torques. The second one permitted to compute the FFS modeled as a force polytope from a musculoskeletal model (MSFP). All the previously mentioned models were compared with a measured force polytope (MFP). To construct it, the maximum isometric forces exerted at the hand were assessed in twenty-six directions of the Cartesian space. Then, the MSFP was applied to the manual wheelchair propulsion in order to characterize the forces applied on the handrim during this task and a new evaluation index of postural performance (IPP) was also introduced.

Three-dimensional, additive printing has emerged as an exciting new technology for the design and manufacture of small spacecraft systems. Using 3-D printed thermoplastic materials, hybrid rocket fuel grains can be printed with nearly any cross-sectional shape, and embedded cavities are easily achieved. Applying this technology to print fuel materials directly into a CubeSat frame results in an efficient, cost-effective alternative to existing CubeSat propulsion systems. Different 3-D printed materials and geometries were evaluated for their performance as propellants and as structural elements. Prototype "thrust columns" with embedded fuel ports were printed from a combination of acrylonitrile utadiene styrene (ABS) and VeroClear, a photopolymer substitute for acrylic. Gaseous oxygen was used as the oxidizer for hot-fire testing of prototype thrusters in ambient and vacuum conditions. Hot-fire testing in ambient and vacuum conditions on nine test articles with a combined total of 25 s burn time demonstrated performance repeatability. Vacuum specific impulse was measured at over 167 s and maximum thrust of individual thrust columns at 9.5 N. The expected ΔV to be provided by the four thrust columns of the consumable structure is approximately 37 m/s. With further development and testing, it is expected that the consumable structure has the potential to provide a much-needed propulsive solution within the CubeSat community with further applications for other small satellites.

Experimental investigations have evaluated the feasibility of using laser-driven plasma microthrusters for small-thrust, high-specific-impulse space maneuvers, particularly for micro- and nanosatellite missions. Recent work made use of the Mach2 hydromagnetics code for the construction of an adequate computational model of the micro-thruster opera- tion. This thesis expounds on this previous work by extending the computational modeling capabilities, allowing for the determination of plasma plume properties and characteristic performance assessment of the microthruster; this allows for further computational investi- gation of the performance improvements achieved by new design considerations. Two par- ticular design changes are implemented and measured: (i) the simulation of microthruster performance intentionally achieving laser-supported detonation of energetic polymer fuels for higher-thrust capabilities, and (ii) the implementation of an axisymmetric nozzle to improve passive solid-fuel performance. The Mach2 hydromagnetics code with the new performance assessment capabilities was used to examine the performance improvement of these new modes of operation; results of the simulations are presented and then evaluated for their use in the overall design of the plasma microthruster. Laser-supported detona- tion shows a tremendous potential increase in the laser momentum coupling coefficient Cm , and demonstrates a much higher thrust; the axisymmetric nozzle varies with nozzle half-angle and length, but still demonstrates expected nozzle trends and improves the laser momentum coupling coefficient, Cm , by up to 230% for some designs considered.

Experimental investigations have evaluated the feasibility of using laser-driven plasma microthrusters for small-thrust, high-specific-impulse space maneuvers, particularly for micro- and nanosatellite missions. Recent work made use of the Mach2 hydromagnetics code for the construction of an adequate computational model of the micro-thruster opera- tion. This thesis expounds on this previous work by extending the computational modeling capabilities, allowing for the determination of plasma plume properties and characteristic performance assessment of the microthruster; this allows for further computational investi- gation of the performance improvements achieved by new design considerations. Two par- ticular design changes are implemented and measured: (i) the simulation of microthruster performance intentionally achieving laser-supported detonation of energetic polymer fuels for higher-thrust capabilities, and (ii) the implementation of an axisymmetric nozzle to improve passive solid-fuel performance. The Mach2 hydromagnetics code with the new performance assessment capabilities was used to examine the performance improvement of these new modes of operation; results of the simulations are presented and then evaluated for their use in the overall design of the plasma microthruster. Laser-supported detona- tion shows a tremendous potential increase in the laser momentum coupling coefficient Cm , and demonstrates a much higher thrust; the axisymmetric nozzle varies with nozzle half-angle and length, but still demonstrates expected nozzle trends and improves the laser momentum coupling coefficient, Cm , by up to 230% for some designs considered.

Le propulseur PEGASES (Plasma Propulsion with Electronegative Gases) est un nouveau type de propulseur électrique pour la propulsion spatiale. Il utilise des ions négatifs et positifs créés par une décharge radiofréquence à couplage inductif pour générer la poussée. L’accélération électrostatique des ions est assurée par un ensemble de grilles polarisées. Un filtre magnétique est utilisé pour augmenter la quantité d'ions négatifs dans la cavité du propulseur. Le propulseur PEGASES est non seulement une source qui permet de créer un plasma d'ions négatifs à forte densité, et même un plasma d'ion-ion, mais il peut également être utilisé comme un propulseur ionique classique. Cela signifie qu'un plasma est créé dans un gaz électropositif (e.g. Xe) et que les ions positifs sont extraits et accélérés. Dans ce cas, il est nécessaire de neutraliser le plasma derrière la zone d'accélération, comme dans d'autres propulseurs ioniques. Les performances du propulseur PEGASES ont été étudiées principalement dans du xénon afin de comparer les résultats obtenus avec les propulseurs ioniques de type RIT. Le propulseur a été étudié à l'aide d'une série de sondes telles qu’une sonde de Langmuir, une sonde plane, une sonde capacitive et un RPA (pour Analyseur à Champ Retardateur). De plus, une sonde en champs croisés ExB a été développée pour mesurer la vitesse des ions quittant le propulseur ainsi que la fraction des différentes espèces ioniques présentes dans le plasma. / The PEGASES thruster (Plasma Propulsion with Electronegative Gases) is a novel type of electric thruster for space propulsion. It uses negative and positive ions produced by an inductively coupled radio frequency discharge to create the thrust by electrostatically accelerating the ions through a set of grids. A magnetic filter is used to increase the amount of negative ions in the cavity of the thruster. The PEGASES thruster is not only a source to create a strongly negative ion plasma or even an ion-ion plasma but it can also be used as a classical ion thruster. This means that a plasma is created and only the positive ions are extracted and accelerated making it necessary to neutralize the plasma behind the acceleration stage like in other ion thrusters. The performances of the PEGASES thruster have been investigated mainly in xenon in order to compare the obtained results with RIT-type ion thrusters. The thruster has been investigated with the help of a variety of probes such as a Langmuir probe, a planar probe, a capacitive probe and a RPA (Retarding Potential Analyzer). In addition, an ExB probe has been developed to measure the velocity of the ions leaving the thruster and to differentiate between the ion species present in the plasma.

<div>When combined with common fuel binders, solid hypergolic fuels can simplify the overall complexity of hybrid rocket systems, as the fuel grain can be ignited and reignited without an external power source or external fluid. In addition, with the hypergolic additive embedded in the binder, the flame zone can be placed at the surface of the grain itself, thereby providing heat to the fuel, improving fuel regression rate, and combustion stability and sustainability. Coupled with high grades of mixed oxides of nitrogen (MON), such hypergolically ignited hybrid configurations are considered a potential propulsion system for a robotic Mars Ascent Vehicle (MAV). Use of the fuel additives and a suitable choice of oxidizer allows for low temperature stability and operation of the propellants, making it an appealing candidate for a simple and storable hybrid propulsion system.</div><div>The first half of this work is dedicated to a very application based study of paraffin based hypergolic hybrids, while the second half of this work, independent from the first, focuses on how theory could help in developing future hypergolic propulsion systems.</div><div>The process undertaken to develop a paraffin based hypergolic hybrid relied heavily on experimental testing of a wide variety of additive loaded fuels with MON to optimize hybrid motor grain parameters with the goals of minimizing ignition delay, improving combustion stability, and promoting sustainment of the flame. MON 3 and MON 25 (3 wt.% or 25 wt.% nitric oxide mixed with nitrogen tetroxide) were used as oxidizers. Through an initial screening process, we selected two solid hypergolic propellants, sodium amide and potassium bis(trimethylsilyl)amide (PBTSA), as additives to promote hypergolic ignition given their low ignition delays with both grades of MON. Iterations on the grain configuration consisted in minimizing the additive loading to simplify the casting process and increase performance, without losing hypergolicity of the grain or hampering combustion sustainability. Using a 90 wt.% hypergolic additive front segment, we were able to light the grain three times using the hypergolic reaction between the additives and MON 3. Once relights achieved, we mainly focused on demonstrating sustained combustion, and determined that, once the front segment depleted, the lack of heat in the system lead the motor to shut down prior to the end of the targeted burn. This led us to add a reactive additive, sodium borohydride, in the main grain, as a way to generate heat in the motor once the front segment was depleted. With the objective of testing relevant conditions for an actual Mars Ascent Vehicle, one of our final tests was done in an altitude chamber, at a 100,000 ft targeted simulated altitude (equivalent to the atmospheric pressure on Mars), with MON 25 as the oxidizer. Using a mixture of sodium amide, PBTSA, and sodium borohydride, we were able to achieve hypergolic ignition in 425 ms (delay to reach 90% of the maximum chamber pressure) at 102,000 ft simulated altitude, for an average chamber pressure of 113 psia.</div><div>During testing we determined that an ideal solid additive should exhibit both low ignition delay with the oxidizer considered, to minimize the motor start up time, and a high heat of combustion, to maximize the energy release and therefore maximize performance. However, the lack of data and theoretical understanding of the reactivity of MONs with non hydrazine based fuels made it challenging to find such an ideal solid additive. Historically, screening processes for new fuel candidates, liquids or solids, have followed a “hit or miss” approach, in which potential fuels were selected based on common characteristics with known hypergols, which is the approach we followed during the development of the hypergolic hybrid. A more robust approach, typically used in the biology and chemistry fields, can be used to predict reactivity of chemicals using statistical analysis. A quantitative structure activity relationship (QSAR) analysis is a statistical analysis used to correlate reactivity to selected molecular descriptors, or properties. Using this approach, one can create models, determined during the QSAR analysis, to predict reactivity of fuel candidates, solely based on their properties. Combined with the recent advancements in computational chemistry and computation of properties, this simple approach has the potential to greatly simplify screening processes for new fuel candidates, as experimental data is not needed anymore. With this method, we were able to fit the logarithmic of the minimum ignition delay for 30 different amines using seven molecular descriptors (heat of formation, heat of vaporization, highest occupied molecular orbital, charge on the nitrogen, rotatable bond count, and ovality), for an R² value of 0.70. While the main motivation behind starting this theoretical work was to optimize for solid additives properties for the hypergolic hybrid configuration described previously, the potential of such model extends to a wider range of propulsion systems (reaction control systems, orbital maneuvering, etc.), and could be expanded to a wider range of oxidizers using machine learning.</div>

No description available.

Colloid thrusters, and electrospray thrusters as a whole, have been around since the 1960s. When they were first developed, the high efficiency and fine thrust control was overshadowed by the high power requirement for such a low thrust that the system provides. This caused the technology to be put on hold for aerospace applications. Now, as small satellites are becoming more prevalent, there has been a resurgence in interest in electrospray thruster technology. The recent advancements in tech- nology allow electrospray thrusters to use significantly less power and occupy less volume than their predecessors. As electrospray technology continues to advance, these thrusters are meeting the demands of small satellite propulsion. As such, in an effort to keep the spacecraft propulsion curriculum current with today’s technology, a colloid thruster is designed, built, tested, and implemented as a laboratory activity at California Polytechnic State University, San Luis Obispo. Electrospray thrusters work by placing a voltage on an ionic liquid and extracting either beads of propellant or ions to generate thrust. By definition, colloid thrusters are a specific class of electrospray thrusters that use solvents, such as glycerol or formamide, to emit droplets or, in special cases, ions to generate thrust. To keep with the University’s “Learn by Doing” pedagogical philosophy, the thruster for this activity is designed to have a tactile and experiential impact on the students. The final design is a scaled up configuration of an existing electrospray design so that the students can easily see each component with the naked eye and can be correlated to a real world thruster that they might see in industry. As a laboratory experiment, the thruster needs to be able to utilize current equip- ment in the Space Environments and Testing Laboratory. One of the Student Vacuum Chambers (SVC) is utilized as well as two 1 kV power supplies and a 100V power supply. An indirect method of measuring performance metrics needs to be developed as there are no thrust balances sensitive enough in the lab designated for undergrad- uate use. As such, the students will be using the mass of the propellant, the time of operation, and knowledge of the propellant’s properties to estimate the performance of the thruster. To prove success of the thruster, a performance profile of the thruster is produced using an indirect method of measurement as well as visual observations of the thruster moving propellant byway of the electrospray theory. The tests show thrusts produced between 96-311 μN with an Isp ranging from 1270-1684 seconds. The visual evidence demonstrates propellant being collected as well as the operation of the thruster under the electrospray theory. The visual evidence also sheds light on which emission mode the thruster is operating at as well as a self-correcting failure mode that was occurring. The thruster is implemented as a lab for Cal Poly’s AERO 402 Spacecraft Propulsion Lab in Fall 2018, and it receives positive feedback from the students through an anonymous survey. While the colloid thruster demonstrates success in meeting performance and pedagog- ical goals, future work should be continued to improve the thruster. Further design and manufacturing work can be undertaken to improve the efficiency and decrease failure due to propellant impingement. Additionally, the procurement of power sup- plies capable of applying higher voltages can provide a greater range of operation which can enable a more dynamic student discovery of electrospray thrusters.

The design process and construction of the initial modular ramjet attachment to the Cal Poly supersonic wind tunnel is presented. The design of a modular inlet, combustor, and nozzle are studied in depth with the intentions of testing in the modular ramjet. The efforts undertaken to characterize the Cal Poly supersonic wind tunnel and the individual component testing of this attachment are also discussed. The data gathered will be used as a base model for future expansion of the ramjet facility and eventual hot fire testing of the initial components. Modularity of the inlet, combustion chamber, and nozzle will allow for easier modification of the initial design and the designs ability to incorporate clear walls will allow for flow and combustion visualization once the performance of the hot flow ramjet is determined. The testing of the blank ramjet duct resulted in an error of less than 10% from predicted results. The duct was also tested with the modular inlet installed and resulted in between a 13-30% error based on the predicted results. Hot flow characteristics of the ramjet were not achieved, and the final cold flow test with the nozzle installed was a failure due to improper configuration of the nozzle. The errors associated with this testing can largely be placed on the poor performance of the Cal Poly supersonic wind tunnel and the alterations made to the testing in an attempt to accommodate these flaws. The final tests were halted for safety concerns and could continue after a thorough safety review.

No description available.