Papers by Francesco Battista
Particle-In-Cell Simulation of Heavy Species in Hall Effect Discharge
Aerotecnica Missili & Spazio, May 20, 2022
AIP Advances, May 1, 2023
Radio-frequency and Helicon Plasma Thrusters have emerged as viable electric propulsion systems d... more 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.
Applied sciences, May 1, 2023
Aerotecnica Missili & Spazio, Apr 1, 2018
Several space electric propulsion devices, such as ion engines and Hall effect thrusters, use hol... more Several space electric propulsion devices, such as ion engines and Hall effect thrusters, use hollow cathodes as the electron sources for providing the necessary electrons for the ionization of the propellant and to neutralize the ion beam leaving the thruster. Since cathode performance is strongly tied with its geometry and size and plasma diagnostics is very difficult to be performed, simple numerical tools are required in order to determine optimal geometries and operative conditions for a given mission profile. The paper presents a preliminary design tool for orificed hollow cathodes. Two physical models compose the tool: a plasma model and a thermal model. A time-independent, volume-averaged model has been developed to determine plasma properties in the emitter and orifice regions. A Lumped Element Thermal Model has been also developed to compute temperature distributions and the respective gradients within the main cathode's element. The study, conducted to validate the results of the models, shows that there is a good agreement with values and trends found in the available literature. The tool is able to estimate performances of new devices by the calculation of cathode working conditions for different geometries given device operating condition and insert material.
Aerospace
The cooling jackets of liquid rocket engines are composed of narrow passages surrounding the thru... more The cooling jackets of liquid rocket engines are composed of narrow passages surrounding the thrust chambers and ensure the reliable operation of the engine. Critical conditions may also be encountered, since the cooling jackets of cryogenic engines, such as those using LOX/LCH4 propellants, are based on a regenerative strategy, where the fuel is used as a refrigerant. Consequently, deterioration modes near where pseudocritical conditions are reached or low heat transfer coefficients where the fuel becomes a vapour and must therefore be managed. The verification of the cooling jacket behaviour to consolidate the design solutions in all the extreme points of the operating box represents a very important phase. The present paper discusses the full characterization of the HYPROB (HYdrocarbon PROpulsion test Bench Program) first unit of the final demonstrator, (DEMO-0A), by considering the working points within the limits of the operating box and comparisons with the nominal conditions ...
Aerospace
Firing test campaigns were carried out on a 200 N thrust-class hybrid rocket engine, using gaseou... more Firing test campaigns were carried out on a 200 N thrust-class hybrid rocket engine, using gaseous oxygen as an oxidizer and a paraffin-wax-based fuel. Different fuel grain lengths were adopted to extend the fuel characterization under different operating conditions, and to evaluate rocket performances and internal ballistics in the different configurations. In addition to data collected under a 220 mm propellant grain length, two further test campaigns were carried out considering 130 mm and 70 mm grain lengths. Two different injector types were adopted in the 130 mm configuration; in particular, a showerhead injection system was used with the aim to contain high-amplitude pressure oscillations observed during some firing tests in this engine configuration. Parameters such as the chamber pressure and temperature inside the graphite nozzle, space-averaged fuel regression rate and nozzle throat diameter were measured. The results allowed for the investigation of different issues rela...
Reduced kinetic mechanism for methane/oxygen rocket engine applications: a reliable and numerically efficient methodology
Combustion Theory and Modelling
Scaling of Magnetic Circuit for Magnetically Shielded Hall Effect Thrusters
Aerotecnica Missili & Spazio
Experimental activities on the paraffin-based fuel MTM in the framework of the PHAEDRA project
AIAA SCITECH 2023 Forum
Numerical Simulations of Fuel Shape Change in Paraffin-Oxygen Hybrid Rocket Engines
Reynolds-averaged Navier–Stokes simulations with submodels of turbulence, chemistry, fluid–surfac... more Reynolds-averaged Navier–Stokes simulations with submodels of turbulence, chemistry, fluid–surface interaction, and radiation are performed in this work to rebuild the internal ballistics of an experimentally tested hybrid rocket engine with paraffin and gaseous oxygen as propellants. Firstly, the effects of the prechamber and postchamber cavities at the initial, average, and final diameter of a reference burn are assessed to be negligible. Then, numerical simulations modeling the fuel shape change in space and time are compared to simulations performed at uniform port radius. The latter provide reasonable regression rate, pressure, and final grain profile predictions with reduced computational cost. On the other hand, the more computationally expensive fuel shape change simulations improve the model prediction capabilities providing a more accurate comparison with experimental data. The fuel shape change approach is finally applied with success to simulations of a throttled burn.
This paper presents the optimization of the zero-dimensional plasma models developed by CIRA for ... more This paper presents the optimization of the zero-dimensional plasma models developed by CIRA for cathodes, obtained removing some of the previous models’ assumptions. Particularly, Xenon electron-neutral cross section has been updated to consider only the elastic collisions: the proposed correlation agrees with the experimental data with an R-squared value of 0.99. The cathode internal total pressure, originally calculated using a theoretical approach, has been estimated by means of an empirical scaling relationship. The formula for the calculation of the internal plasma resistance has been revisited. A plasma model for the keeper region has been proposed, too. A LaB6 based emitter OHC fed by Xenon has been selected as baseline geometry to test the model. The trends of electron temperature, plasma densities, neutral densities and wall temperatures, as function of mass flow rate and discharge current, are in agreement with those reported in literature, as well as their distributions ...
Development of a cathode for low-power Hall Effect Thruster
Two-Hundred-Newton Laboratory-Scale Hybrid Rocket Testing for Paraffin Fuel-Performance Characterization
Journal of Propulsion and Power, 2018
A series of firing tests have been performed on a laboratory-scale hybrid rocket engine of 200 N ... more A series of firing tests have been performed on a laboratory-scale hybrid rocket engine of 200 N class, fed with gaseous oxygen through a converging nozzle injector, to assess the mechanical feasib...
Thermostructural Analyses Supporting the Design of the HYPROB Heat Sink Subscale Breadboard
Volume 1: Advances in Aerospace Technology, 2014
Today’s rocket engines regeneratively cooled using high energy cryogenic propellants (e.g. LOX an... more Today’s rocket engines regeneratively cooled using high energy cryogenic propellants (e.g. LOX and LH2, LOX and LCH4) play a major role due to the high combustion enthalpy (10–13.4 kJ/kg) and the high specific impulse of these propellants. In the frame of the HYPROB/Bread project, whose main goal is to design build and test a 30 kN regeneratively cooled thrust chamber, a breadboard has been conceived in order to:• investigate the behavior of the injector that will be employed in the full scale final demonstrator,• to obtain a first estimate of the heat flux on the combustion chamber for models validation,• to implement a “battleship” chamber for a first verification of the stability of the combustionThe breadboard is called HS (Heat Sink) and it is made of CuCrZr (Copper Chromium Zirconium alloy), Inconel 718 and TZM (Titanium Zirconium Molybdenum alloy). The aim of the present paper is to illustrate the thermostructural design conducted on the breadboard by means of a Finite Element Method code taking into account the viscoplastic behavior of the adopted materials. An optimization process has been carried out in order to keep the structural integrity of the breadboard maximizing the life cycles of the component. Heat fluxes generated by combustion gases have been evaluated by means of CFD quick analyses, while convection and radiation with the external environment have not been considered in order to be as conservative as possible from a thermostructural point of view. Transient thermal analyses and static structural analyses have been performed by means of ANSYS code adopting an axisymmetric model of the chamber. These analyses have demonstrated that the Breadboard can withstand the design goal of 3 thermo-mechanical cycles with a safety factor equal to 4 considering a firing time equal to 3 seconds.Copyright © 2014 by ASME
MHD Experiment at CIRA GHIBLI Plasma Wind Tunnel
Oxygen–methane rocket thrust chambers: Review of heat transfer experimental studies
Acta Astronautica, Aug 1, 2023
High Power Hall Thrusters Design Options IEPC-2007-311
The development of an analytical scaling model coupled with a statistical database defines the op... more The development of an analytical scaling model coupled with a statistical database defines the opportunity of a wide and more flexible way of scaling Hall thruster in order to follow different design requirements. This paper is focused on the scaling methodology applied to a high power Hall thruster and it aims to provide different design options for this class of thrusters. In order to obtain a scaling model as effective and rational as possible, a vector of fundamental parameters, defining the thruster geometry and performances, has been created. Three geometric parameters has been chosen: the channel length, the channel width and the channel average diameter. The other two fundamental parameters are the gas number density in the injection plane and the applied discharge voltage. Given a set of basic physical relations which link these five parameters to all the other thruster performance, it is possible to build a scaling matrix. This matrix is a useful tool, which can then be us...
Validation of Conjugate Heat Transfer Model for Rocket Cooling with Supercritical Methane
Journal of Propulsion and Power, 2016
A numerical solver able to describe a rocket engine cooling channel fed with supercritical methan... more A numerical solver able to describe a rocket engine cooling channel fed with supercritical methane is validated against experimental data coming from a test article conceived and tested by the Italian Aerospace Research Center. The multidimensional conjugate heat transfer model numerically solves the Reynolds-averaged Navier–Stokes equations for the coolant flow and the Fourier’s law of conduction for the heat transfer within the wall. In this study, an experimental test case is reproduced in detail in order to evaluate the influence of partially unknown parameters, such as surface roughness and wall thermal conductivity, and of operative parameter uncertainty, such as the coolant mass flow rate and input heat transfer rate. The comparison made with respect to the wall temperature and coolant pressure drop of the whole set of experimental data provides complementary information that allows better understanding of experiments and infers possible deviations from the expected behavior.
Characterization of a new Mach 9 nozzle for the HEAT hypersonic wind tunnel
ABSTRACT A new Mach 9 contoured nozzle to use with air was designed and realized at Alta SpA with... more ABSTRACT A new Mach 9 contoured nozzle to use with air was designed and realized at Alta SpA with the aim to produce a uniform core flow with a diameter of at least 80 mm. The design was iteratively carried out using engineering codes and CFD simulations by CIRA. The characterization activity was carried out mapping the complete test section in terms of pitot pressure and total enthalpy and measuring the pressure and heat flux distribution on the nozzle internal walls. The flow before the convergent was characterized by means of total pressure measurements and spectroscopy. A numerical rebuilding of the test was performed by CIRA and PoliTO and compared with experimental data. The paper will briefly describe the design phase and will present all the characterization results.
Kinetic assessment of high pressure methane-oxygen ignition and combustion was performed as preli... more Kinetic assessment of high pressure methane-oxygen ignition and combustion was performed as preliminary step towards Liquid Rocket Engine (LRE) on-ground technological demonstrator design, manufacturing, development, experimentation and optimization. Ignition delay times were calculated using a commercial zero-dimensional software (ANSYS ChemkinPro 17.0) and literature available both detailed and reduced kinetic schemes of methane oxidation at temperatures, pressures and mixture ratios generally experienced in LRE combustion chamber. Comparison of kinetic parameters, among the various investigated kinetic schemes, allows the identification of the most suitable mechanism for analysing, by means of a computational approach, the reacting flow dynamics of methane-oxygen mixtures within a high pressure LRE and build up a numerical method for improving the predictive capacity of Computational Fluid Dynamics (CFD) simulations used for design criteria. Introduction At the moment, the access...
Uploads
Papers by Francesco Battista