Hall Effect Thrusters (HETs) are nowadays widely used for satellite applications because of their efficiency and robustness compared to other electric propulsion devices. Computational modelling of plasma in HETs is interesting for several reasons: it can be used to predict thrusters’ operative life; moreover, it provides a better understanding of the physical behaviour of this device and can be used to optimize the next generation of thrusters. In this work, the discharge within the accelerating channel and near-plume of HETs has been modelled by means of an axisymmetric hybrid approach: a set of fluid equations for electrons has been solved to get electron temperatures, plasma potential and the discharge current, whereas a Particle-In-Cell (PIC) sub-model has been developed to capture the behaviour of neutrals and ions. A two-region electron mobility model has been incorporated. It includes electron–neutral/ion collisions and uses empirical constants, that vary as a continuous function of axial coordinates, to take into account electron–wall collisions and Bohm diffusion/SEE effects. An SPT-100 thruster has been selected for the verification of the model because of the availability of reliable numerical and experimental data. The results of the presented simulations show that the code is able to describe plasma discharge reproducing, with consistency, the physics within the accelerating channel of HETs. A small discrepancy in the experimental magnitude of ions’ expansion, due probably to boundary condition effects, has been found.

Natural convection heat transfer is found in many industrial applications such as nuclear technologies, electronic circuit board cooling, solar panel cooling and many other fields. When natural convection heat transfer coefficients are insufficient, passive heat transfer enhancement devices (called ribs/fins) are often used. In this paper, the effect of periodic patterns of protrusions (ribs) on the free-convection heat transfer of a vertical plate, with a uniform heat flux rate boundary condition, is experimentally investigated. The convective fluid is air. Two-dimensional, high resolution heat transfer measurements are performed by using infrared thermography and the heated thin foil technique. Experiments are performed on two types of ribs pattern topology: single or two staggered rows of ribs inclined at different angles and single or two-staggered rows of V-ribs

Radio-frequency and Helicon Plasma Thrusters have emerged as viable electric propulsion systems due to their high plasma density, thrust density, and useful life. Helicon Plasma Thruster (HPT) is a very attractive technology because it could use many propellants and does not require hollow cathodes or grids, overcoming their associated critical erosion problem and extending the thruster’s lifetime to some tens of thousands of hours. Despite the fact that high-power HPTs have reached 30% efficiency in laboratory configurations, sophisticated numerical models are required for a deeper understanding of the main plasma phenomena and for the preliminary design to increase the very low HPT’s efficiency (3–7%) typical of the low-power class thrusters. The paper focuses on the development of a model for the low-medium power range (50–2000 W) of HPTs design. Starting from Lafleur’s model, it has been improved with the hypothesis of neutral gas being expelled at the real thruster’s wall operative temperature (300–600 K) in place of the more frequent laboratory temperature assumption (300 K). This hypothesis affects total thrust and specific impulse by about 10%. A parametric analysis of the slenderness ratio (chamber length-to-radius) has been conducted. The results showed that slender configurations lead to higher efficiencies. Downstream from the numerical model validation, a tool for the global design has been built with the Particle Swarm Optimization (PSO) technique that leads to optimal thruster configuration. This tool has been used to design a 4 mN HPT tuning the PSO in order to minimize the dimensions and the weight according to the assigned mission constraints (i.e., power, thrust, and weight). A total efficiency of 10.4% results.

Solid Rocket Motors frequently experience unsteady gas motions and combustion instabilities. Pressure oscillations are a well-known problem of large solid rocket motors (e.g. those of the US Space Shuttle, Arianne 5 P230 and P80). Pressure oscillations lead to thrust oscillations which can generate unstable dynamic environments for the rest of the launcher up to the payload. This kind of instability is governed by the flow behavior of the combusted gas combined with pressure fluctuations and acoustic resonances within the combustion chamber. In the present investigation a computational analysis of the combusted gases passing through the chamber of such a solid rocket motor has been conducted, with particular attention to Corner and Parietal Vortex Shedding instabilities, inside the core section of the motor together with a study of the associated pressure oscillations. Nomenclature Downloaded by Mario Panelli on July 18, 2017 | http://arc.aiaa.org |

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