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The paper describes the activities of the IRENE (Italian Re-Entry NacellE) program, funded by the Italian Space Agency (ASI) and the European Space Agency (ESA), which are aimed at realizing an innovative deployable (umbrella-like) heat shield concept developed by ALI, CIRA and University of Naples Federico II. IRENE is a capsule whose innovative characteristics are the heat-shield opening mechanism (umbrella-like configuration) and the (off-the-shelf) materials used for the thermal protection, both covered by international patents.The “umbrella-like” deployable heat shield reduces the capsule ballistic coefficient, leading to acceptable heat fluxes, mechanical loads and final descent velocity. The feasibility study of IRENE has been carried out in 2011. The Thermal Protection System (TPS) materials have been tested in the SPES hypersonic wind tunnel at the University of Naples, and in the SCIROCCO Plasma Wind Tunnel at CIRA in 2011. Thanks to the promising results of the first activities, ASI and ESA decided to fund, in the frame of GSTP program, the current phase of the program called MIFE “MINI-IRENE Flight Experiment”. To test the functionality of the MINI IRENE deployable heat shield for the atmospheric re-entry, two different-scale demonstrators have been designed and manufactured, achieving an overall TRL 6. A Ground Demonstrator (GD), representative of the TPS materials (nose and flexible part) and of the mechanisms of the FD, has been successfully qualified in SCIROCCO (Plasma Wind Tunnel) in 2018. A Flight Demonstrator (FD), designed to sustain the dynamic pressure and mechanical loads encountered during launch and microgravity phases of the re-entry mission, has been subjected to a space qualification campaign between the end of 2019 and the beginning of 2020. The Flight Demonstrator will be included as a secondary payload in the interstage adapter of a VSB-30 Sounding Rocket that will be launched in June 2022, from the ESRANGE (Sweden) launch base. Then, during the ascent phase of the payload section, it will be ejected, perform a 15-minutes ballistic flight, re-enter the atmosphere and hit the ground. As a further application of “IRENE technology”, ALI has proposed the “Small Mission to Mars” (SMS) project. SMS is a low-cost system suitable for the entrance in the atmosphere and the operations on the ground of Mars. The SMS feasibility study was funded by the European Space Agency (ESA) in 2016. In the paper are also illustrated the results of the feasibility study of SMS, including a description of the mission profile, launch and escape phases, interplanetary trajectory, Mars approach, entry, descent and landing (EDL), and payload deployment and operations.

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 paper presents an extensive experimental campaign over the effect of SiC concentration on the aero-thermal behavior of ultra-high temperature ceramics in the hypersonic atmospheric re-entry environment. Four compositions made of ZrB2 with different amount of SiC from 5 to 18 vol.% were exposed to the supersonic plasma flow of an arc-heated plasma wind tunnel, at specific total enthalpies up to 20 MJ/kg measuring their surface temperature by non-intrusive diagnostic equipment, including two-color pyrometers and an infrared thermo-camera. As SiC content increases, the maximum steady-state temperature reached on the surface decreased and emissivity value are higher. During some tests, a spontaneous temperature jump in the order of 400 K was observed, which only occurred on the front surface of the sample. Surface temperatures over 2800 K were measured after a temperature jump. The composition which experienced the jump showed an external surface reaction made of only zirconia layer on the front surface, probably appearing upon complete removal of liquid borosilicate glass which forms during exposure to oxygen and is still present on the surface of samples with higher SiC content. The experiments demonstrated that the temperature jump, closely related to the overall thermal stability of the material, appears favoured by lower SiC amount, but it can be triggered also in case of larger SiC content (up to at least 15% vol.), as long as the flow total enthalpy and the exposure time are sufficiently high.

In the framework of the Horizon 2020 project C3HARME, an experimental campaign has been carried out to characterize a new class of Ultra-High-Temperature Ceramic Matrix Composites for near-zero ablation Aerospace Thermal Protection Systems. Small sized specimens, with ZrB2-based matrix and different carbon fiber architectures, were exposed to a supersonic flow of simulated air generated by an arc-jet wind tunnel, achieving specific total enthalpies up to 20 MJ/kg, 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 a good performance of the analyzed samples, although with some mechanical resistance issues. Surface temperatures were monitored by means of infrared pyrometers and a thermo-camera, and during most of the tests a spontaneous temperature jump was observed, with temperatures that reached values over 2800 K at the steady state. Computational Fluid Dynamics simulations allowed for the rebuilding of the thermo-fluid-dynamic and chemical flow field. Moreover, it was possible to propose a 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.

MINI-IRENE is the Flight Demonstrator (FD) of IRENE, a new-concept capsule with a variable geometry, originally conceived by ASI to widen the range of available platforms to retrieve payloads and/or data from low Earth orbit. The main characteristics of IRENE is the “umbrella-like" deployable front structure that reduces the capsule ballistic coefficient, leading to acceptable heat fluxes, mechanical loads, stability and final descent velocity. The feasibility study of the IRENE deployable re-entry system has been carried out in 2011. The TPS materials, selected for the nose cone and for the flexible umbrella shield, have preliminarily been tested in the SPES hypersonic wind tunnel at the University of Naples, and in the SCIROCCO Plasma Wind Tunnel at CIRA. Such successful tests and the preliminary experimental results proved the concept feasibility and the viability of commercial materials for lowcost re-entry nacelles. After the preparation phases A and B were successfully completed, the European Space Agency funded the current phase of the program. The object is to design and built a Flight Demonstrator and a Ground Demonstrator to prove, with a suborbital flight and with a Plasma Wind Tunnel (PWT) test campaign, the functionality of the deployable heat shield. The Flight Demonstrator shall be included as a secondary payload in the interstage adapter of a VSB-30 launcher from ESRANGE. It shall then be ejected during the ascent phase of the payload section, after its separation from the booster at an altitude of about 150 km, perform a 15 minutes ballistic flight, re-enter the atmosphere and hit the ground. The Ground Demonstrator, representative of the Thermal Protection System of the Flight Demonstrator, shall be instead exposed to a heat flux similar to that expected for an atmospheric re-entry from low Earth orbit inside the SCIROCCO Plasma Wind Tunnel at CIRA. The paper, after a short description of the mission profile both for orbital and suborbital flights, focuses on the aerodynamics and flight mechanics activities together with an overview of the design of the mechanism that will deploy the umbrella and that will cope with the re-entry environment.

In the framework of the Horizon 2020 project C3HARME, an experimental campaign has been carried out to characterize a new class of Ultra-High-Temperature Ceramic Matrix Composites (UHTCMC) for near-zero erosion rocket nozzles. A novel test set-up has been developed to the purpose of testing small-sized specimens, with ZrB2- based matrix and carbon fibers, exposing them to the supersonic exhaust plume of a 200N-class hybrid rocket nozzle, employing gaseous oxygen as oxidizer and High-Density PolyEthylene as fuel. The aim of the tests was to reproduce realistic rocket nozzles operating conditions, in order to demonstrate the ability of the specimens to preserve their functional integrity in a relevant environment. After that, a UHTCMC nozzle throat insert has been manufactured and experimentally tested to verify the erosion resistance and evaluate the effects on the rocket performance by comparison with those obtained in similar operating conditions employing a graphite nozzle. Computational Fluid Dynamics simulations supported the experimental activities allowing for the rebuilding of the thermo-fluid-dynamic and chemical flow field and the characterization of the test conditions.&nbsp

In space propulsion hydrogen peroxide is a “green” solution that overcomes the problem of toxicity related to hydrazine-based systems. This work aims at designing, developing, and characterizing a monopropellant thruster for small satellites, based on hydrogen peroxide. The engine is intended for use on a 12-Unit CubeSat in Low Earth Orbit. Based on mission requirements a catalytic chamber and a nozzle were designed and a ground-test breadboard was manufactured. The prototype was assembled and tested on ground and measured engine performances are discussed highlighting their dependence on the nozzle throat diameter, that primarily affects the chamber pressure. Finally, effects of the mass flow rate on the system reactivity and steady state conditions are also investigated.

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.

Qualification and reusability campaigns were performed on ultra-high temperature ceramic matrix composites (UHTCMCs) made of a ZrB2-SiC matrix with short/long carbon fibre to assess their performance as thermal protection systems. Heat fluxes and stagnation pressures were set following those of reference re-entry missions. An interdisciplinary analysis followed by combining infrared measurements of the thermal response, estimation of surface properties and microstructural characterization. Compared to conventional C/SiC composites, the increased erosion resistance of the new UHTCMCs was assessed in ultra-hot regime: although these experienced much higher thermal loads, they underwent minimal erosion and preserved their structural stability even after multiple exposures.

This work presents a novel approach for the modeling of the entrainment in the numerical simulation of the internal ballistics of hybrid rocket engines with paraffin-based fuels. This model, coupled with a more sophisticated gas–surface interaction treatment, is an improvement of the model previously developed by some of the authors, which was based on some oversimplifying assumptions. Indeed, the old entrainment model was performed in a close range of averaged oxidizer mass flux, and it made the overall numerical model not scalable on different motor sizes. Therefore, a new correlation is introduced, which is based on the Reynolds number and takes into account the dependence of the entrained fraction on the shear stress exerted by the gas flow and the tube diameter. Firstly, the new and the old numerical models are compared in order to highlight the improvement obtained by the current efforts. Then, the model has been validated on experimental tests involving two different thrust class motors. Finally, the effect of the motor size on the fuel consumption is shown, thus revealing the crucial influence of the recirculating zone extension due to the oxidizer axial injection.

An increasing number of institutions and companies is recently showing increasing interest in supersonic and hypersonic flight with particular focus on sub-orbital flight and high-speed point-to-point transportation. In this scenario, the University of Naples “Federico II” is currently studying advanced concepts of supersonic/hypersonic aircraft and suitable high-speed propulsion system integration for efficient travel in the Mach 4-5 speed range. Within this framework, a collaboration is in progress with the SME Blue Engineering. In the present work a study is undertaken to investigate the performances along with aerodynamic design of the supersonic intake for the turboramjet engines and the engine-airframe integration for a Mach 4 supersonic business jet. Firstly, attention is focused on directional stability and trimmability enhancement acting on aircraft configuration. The following section is dedicated to the aerodynamic design of a high-efficiency intake adapted to different flight conditions. To conclude, the last section addresses the sensitive issue of airframe-engine integration. The configuration analysis has been carried out with the quick and low-cost software Missile DATCOM, well suited for aircraft preliminary design. For the supersonic intake aerodynamic design and the airframe-engine integration analysis, the commercial Navier-Stokes solver Siemens STAR CCM+ has been used. An improved configuration has been proposed, actioning on vertical tail and wing for the directional stability enhancement, while the trim conditions have been improved using a canard. As for the supersonic intake aerodynamic design, the choice fell on mixed-compression variable-geometry. In particular, a total spike forward translation of 0.65 m guarantees adaptive operations along the supersonic trajectory. In addition, the choice of mixed-compression configuration ensured high efficiency values thanks to the multiple shock waves’ reflection in the supersonic diffuser. For what concerns the engine-airframe integration, the engine location above the wing has been shown to be not suitable due to the strong aerodynamic interference that reflects in a completely off-design intake functioning. Therefore, the configuration has been upgraded integrating the engine with the wing avoiding the issues experienced in the previous case

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.

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.

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.

MINI-IRENE is the Flight Demonstrator of IRENE, a new-concept capsule with a variable geometry, originally conceived by ASI to widen the range of available platforms to retrieve payloads and/or data from low Earth orbit. The main characteristics of IRENE is the “umbrella-like" deployable front structure that reduces the capsule ballistic coefficient, leading to acceptable heat fluxes, mechanical loads, stability and final descent velocity. Following the feasibility studies carried out since 2011, with also preliminary Thermal Protection System materials tests in plasma wind tunnels, the objective is now to design and build a Flight Demonstrator and a Ground Demonstrator to prove, with a suborbital flight and with a Plasma Wind Tunnel (PWT) test campaign, the functionality of the deployable heat shield. The Flight Demonstrator shall be included as a secondary payload in the interstage adapter of a VSB-30 launcher from ESRANGE, then ejected during the ascent phase of the payload section, perform a 15-min ballistic flight, re-enter the atmosphere and hit the ground. The Ground Demonstrator, representative of the Thermal Protection System of the Flight Demonstrator, shall be instead exposed to a heat flux similar to that expected for an atmospheric re-entry from low Earth orbit inside the SCIROCCO Plasma Wind Tunnel at CIRA. The paper, after a short description of the mission profile both for orbital and suborbital flights, focuses on the aerodynamics and flight mechanics activities held for the suborbital flight and PWT test campaigns.

This work is focused on the aerodynamic analysis of a small satellite provided with a deployable aero-brake. The satellite is intended to perform a completely aerodynamic de-orbiting maneuver from Low-Earth-Orbit. A brief discussion about the aerodynamic effects of the position of the aero-brake along the longitudinal axis of a simplified axisymmetric system is presented. Moreover, a more complex architecture, envisaging deployable solar panels for the enhancement of power generation along the orbital path, is proposed and analyzed. The present paper is aimed at the evaluation of the influence of such a configuration on the satellite aerodynamic parameters. Computations have been carried out by means of a Direct Simulation Monte Carlo (DSMC) code at altitude of 150 km, velocity of 7800 m/s and in the interval of angle of attack 0–180 deg with a spacing of 10 deg. The results verified that the deployable solar panels strongly influence Aerodynamics of the satellite. One of the most relevant aspects is the variation of the longitudinal stability equilibrium that becomes more stable. Furthermore, the deployable solar panels increase the aerodynamic drag when the aero-brake is closed, affecting the drag modulation capability.

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.