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Activities carried out by the Italian Aerospace Research Centre, concerning studies about hybrid rocket engine technology with paraffin-based fuels, are described in this paper. The work has been performed in the framework of the HYPROB project, Demonstrators Line. Preparatory activities, including experimental firing tests on a subscale breadboard of 200 N thrust class, are first described. These tests allowed evaluating the paraffin-based grain characteristics, in terms of regression rate and mechanical properties, and validating numerical models adopted to support the demonstrator design. The hybrid demonstrator, which is a 1000 N thrust class, is later described in detail. Main results of the analyses, carried out with engineering tools and numerical codes, are presented and discussed. The status of the manufacturing activities and future works are finally reported.

This work deals with experimental results obtained during the 67th Parabolic Flight Campaign of the European Space Agency. The capillary flow of self-rewetting fluids i.e. dilute aqueous solutions of long chain alcohols with an unusual surface tension behaviour, has been investigated onboard a zero-g plane under different gravity levels. Such mixtures have been extensively investigated on ground as working fluids for two-phase heat transfer devices. The presence of small proportions of alcohols in water changes both the wetting and surface tension properties of the mixture. Contrary to ordinary liquids, the surface tension becomes an increasing function with temperature that, in addition to the variation induced by the preferential evaporation of the more volatile component, provides a reverse Marangoni flow along the liquid-vapor interfaces driven towards the hotter regions. As working fluids for heat pipe systems, self-rewetting fluids show better properties, i.e. lower thermal resistance, enhanced dry-out limit and more stable behaviour. The parabolic flight experimental configuration includes a V-shaped groove channel partially filled with a water/butanol mixture and equipped with a top transparent window and a lighting system, enabling visualization of the liquid in the groove with a CCD camera. Results show that the liquid film distribution is affected from the gravity levels. During the parabolic manoeuvres, the liquid remains confined inside the groove channel and increasing the power level the thickness gradually decreases. The results are explained with respect to the thermo-physical properties of the self-rewetting mixture and discussed in relation to the experiments carried out in normal gravity condition.

The aim of the present paper is the investigation of a monopropellant engine based on the exothermic decomposition of nitrous oxide, as a “green” substitute of the most common but highly toxic hydrazine-based systems. First, a general procedure for the design of such a system, in terms of sizing of the catalytic chamber and of the nozzle for prescribed mission requirements, is defined. This procedure is applied for the design of an 800-mN-class thruster prototype, with catalysts based on palladium as active phase on alumina pellets support. Such prototype is experimentally tested highlighting first the nitrous oxide decomposition behavior and issues related to an optimization of system thermal insulation for reducing energy losses. Finally, the measured engine performance are presented and discussed, including a comparison with the results obtained with the same system in a cold-flow operation mode, showing the gain in terms of increased specific impulse and reduced propellant consumption due to the exploitation of the thermal energy obtained from nitrous oxide decomposition.

An experimental study on the catalytic decomposition of the nitrous oxide with catalysts based on palladium as active phase on alumina pellets support has been carried out and is presented in this paper, with the aim of inquiring the potentialities of nitrous oxide for space propulsion applications, especially as monopropellant for low-thrust engines. Different tests were carried out, with different system configurations and varying the propellant mass flow rate, to investigate the main dependences of the decomposition efficiency with the characteristic parameters of the system. The results showed that an optimal operating condition exists in terms of the catalytic bed load and the gas hourly space velocity. Finally, a procedure was designed for increasing the catalytic system initial temperature, and a new experimental campaign was carried out, highlighting the effect of the latter parameter on the nitrous oxide decomposition behavior.

This paper presents the results of experimental tests for the characterization of Ultra-High-Temperature Ceramic Matrix Composite (UHTCMC) materials for near-zero erosion rocket nozzles. Two dedicated test set-ups were developed for preliminary screening of material candidates in a representative environment, characterized by relevant heat flux and temperature. The experimental set-up was based on a lab-scale 200N-class hybrid rocket engine, employing gaseous oxygen as the oxidizer and High-Density PolyEthylene as fuel; the configurations included free-jet test, in which small button-like samples were exposed to the supersonic exhaust jet of the rocket nozzle; and chamber inserts, in the shape and size of an annular element, placed inside the rocket combustion chamber. Computational Fluid Dynamic simulations, for modeling heat transfer and combustion chemical reactions, complemented the experimental observations and supported the characterization of test conditions. Samples with ZrB2-SiC matrix and continuous or chopped carbon fibers, sintered by either Hot Pressing or Spark Plasma Sintering were tested. Free-jet test samples demonstrated a substantially improved erosion resistance with respect to conventional graphite and in one case a negligible material recession. UHTCMC samples erosion was associated to the occurrence of a rapid rise in surface temperature, which achieved values over 2900 K. Chamber inserts, besides confirming the outstanding erosion resistance of UHTCMCs with respect to traditional materials (i.e. C/SiC), proved that long-fibers samples with sufficient porosity are more likely to withstand thermal shocks typical of the rocket combustion environment.

Two samples of Ultra-High-Temperature Ceramic Matrix Composites, with carbon fibers in a ZrB2-SiC matrix, were exposed to supersonic dissociated air flow, simulating the atmospheric re-entry environment, in an arc-heated facility at specific total enthalpies up to 20 MJ/kg. Surface temperatures, exceeding 2400 K, were monitored by non-intrusive infrared equipment, which allowed detecting thermo-chemical surface instability phenomena. A zirconium oxide layer formed on the surface, below which the original material is perfectly preserved. Numerical simulations allowed describing the flow field around the samples and characterizing the materials behavior, in terms of thermal conductivity, catalycity and oxidation effects at high enthalpies.

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 feasibility and regression rate of a newly developed paraffin-based fuel. Such an injector configuration, by producing recirculation at the motor head hand, has been already demonstrated to influence the standard fuels regression rate, which yields an increase with the port diameter at given mass flux. In this study, paraffin-fuel regression rate dependence on the mass flux and grain port diameter in the form of a power function is determined to be similar to that established with polymeric fuels, despite the different mechanism of consumption that involves the fuel surface liquid-layer instability other than the vaporization typical of classical polymers. Comparison with some data in the literature is presented. Data retrieved from the testing campaign are compared with numerical results obtained by adopting a simple but efficient modeling strategy and a commercial solver. The numerical solution gives evidence of the recirculating flow at the injector exit, which is also responsible for the paraffin contamination observed in the motor prechamber. A good agreement is found with chamber pressure experimentally measured.

The computational fluid dynamics of hybrid rocket internal ballistics is becoming a key tool for reducing the engine operation uncertainties and development cost as well as for improving experimental data analysis. Nevertheless, its application still presents numerous challenges for the complexity of modeling the phenomena involved in the fuel consumption mechanism and its coupling with the chemically reacting flowfield. This paper presents a review of the computational thermo-fluid-dynamic models developed for the internal ballistics of hybrid rockets burning gaseous oxygen with classical polymeric or paraffin-based fuels, with a special focus on the interaction between the fluid and the solid fuel surface. With the purpose of predicting the local fuel regression rate, which is the main parameter needed for the hybrid rocket design, the model is coupled with an improved gas/surface interface treatment based on local mass, energy and mean mixture-fraction balances, combined to either a pyrolysis-rate equation in the case of classical polymers, or to an additional equation for the liquid paraffin entrainment fraction of the total fuel consumption rate. A number of experimental test cases obtained from the static firing of two different laboratory-scale rockets are simulated to determine the models’ capabilities, showing very good agreement between the calculated and measured fuel regression rates with both standard pyrolyzing and liquefying fuels. The prediction of the chamber pressure measured with paraffin fuel resulted in it being more cumbersome for the single-phase flow assumption. The advantages and limitations of the models are discussed.

The present paper describes the design and testing activities carried out on a 1000 N hybrid rocket demonstrator, including a preliminary data assessment. Gaseous oxygen and a paraffin-based solid fuel represent the hybrid propellants. The demonstrator configuration and the design logic are presented first. Later, the experimental results of the test campaign are reported. Results include numerous acquisitions, such as chamber pressure, temperatures and thrust. A preliminary assessment of the experimental data is finally discussed. The demonstrator provided a stable combustion in all the testing conditions and performances in line with the expectations. Throttling capability of the test article was also demonstrated.

With the aim of characterizing the behaviour of paraffin-based fuels for hybrid rocket application, research activities are currently ongoing at University of Naples “Federico II” by means of experimental testing and numerical modelling. In the present work, first the results of an experimental campaign performed on a laboratory-scale hybrid rocket engine of 200 N class, with gaseous oxygen and paraffin-fuel grains, are presented. The relatively vast quantity of the collected data supported the definition, tuning and validation of proper computational fluid dynamic numerical models for the prediction of the fuel consumption behaviour. After that, effects of grain geometrical dimensions on the fuel regression rate are investigated, starting from preliminary considerations obtained for the case of a classical polymeric fuel and then extending the study to the case of paraffin-based fuel.

A computational thermo-fluid-dynamic model for the simulation of the internal ballistics of hybrid rockets burning gaseous oxygen and paraffin-based fuel is presented in the present work. The main objective is the prediction of the solid fuel regression rate, which is calculated with an improved gas/surface interface treatment based on local mass, energy and mean mixture-fraction balances and an additional equation for modelling the liquid droplets entrainment contribution to the total fuel mass flow rate. Parametric analyses have been carried out to assess the effect of fuel physical properties on the results. Comparison between numerically calculated and experimentally measured regression rate axial profiles retrieved from three firing tests performed with two laboratory scales hybrid rocket motors are outlined to address preliminary validation of the model and identify possible future improvements.

An experimental campaign has been carried out to characterize a new class of Ultra-High-Temperature Ceramic Matrix Composites for near-zero ablation Thermal Protection Systems. Small-sized specimens, with ZrB2-based matrix and different carbon fiber architectures, were exposed to a simulated air supersonic flow generated by an arc-jet wind tunnel, achieving specific total enthalpies up to 20 MJ/kg and cold wall fully catalytic heat fluxes over 5 MW/m2, in an aero-thermo-chemical environment representative of atmospheric re-entry. Ablation rates were estimated by means of mass and thickness measurements before and after testing, demonstrating an excellent performance of the developed materials. Surface temperatures were monitored by means of infrared pyrometers and a thermo-camera, and during all the tests a spontaneous temperature jump was observed, with temperatures that reached values over 2800 K at the steady state. Post-test microstructural analyses revealed the formation of a porous oxide layer with a thickness of few hundred microns, mainly consisting of zirconia, with substantial removal of both SiC and carbon fibers. Below the oxide, the bulk material was unaffected. Computational Fluid Dynamics simulations allowed rebuilding the thermo-fluid-dynamic and chemical flow field. Moreover, it was possible to propose an innovative correlation of the temperature jump with an increased catalytic activity and a dramatic reduction of the thermal conductivity of the oxide layers forming on the exposed part of the sample, which anyway had a key role in preserving the unoxidized bulk materials at reasonable temperatures.