Aero-thermo-dynamic Study on Thermal Protection Systems and Rocket Nozzles in Innovative Ceramic Materials

New-generation hypersonic and reusable re-entry vehicles set increasingly demanding requirements for the development of high-performance Thermal Protection Systems (TPS), due to the challenges of extremely harsh aero-thermo-dynamic conditions characteristic of atmospheric re-entry, including hypersonic Mach numbers, temperatures above 2000°C, the activation of gas dissociation/recombination reactions at extremely low oxygen partial pressures, which can substantially enhance the heat flux on the exposed surface of the spacecraft. On the other hand, challenges for solid and hybrid rocket technologies include the design and fabrication of non-eroding firing thrusters able to survive severe thermal-structural and thermal-chemical combustion environments without cooling systems. The inner surface of the exhaust nozzle, through which the propellant flow is accelerated to supersonic conditions, is very critical in this sense, as it is subjected to the highest shear stresses, pressures and heat fluxes in a chemically aggressive environment. These severe conditions usually lead to removal of surface material (ablation) due to heterogeneous reactions between oxidizing species in the hot gas and the solid wall. Because of the material erosion, there is an enlargement of the nozzle throat section and a consequent decrease of the rocket thrust, with detrimental effects over the motor operation. Thus, the requirement that dimensional stability of the nozzle throat should be maintained makes the selection of suitable rocket nozzle materials extremely hard. Over the last decades, research identified Ultra-High-Temperature Ceramic (UHTC) materials, based on transition metals carbides and diborides, as potentially promising candidates for these applications, especially in light of their high melting temperatures, strength and ablation resistance at temperatures over 2000°C. Nevertheless, some issues related to poor oxidative behaviour and mechanical properties (damage tolerance, fracture toughness, thermal shock resistance) of single and multi-phase UHTCs at high temperatures limit the applicability of these materials. The introduction of SiC or other silicon based ceramics as minority phase, in the form of particles, short/long fibres or whiskers, in the main refractory ceramic has been often proposed to improve damage tolerance and oxidation resistance at intermediate temperature, essentially thanks to the formation of a low-viscosity borosilicate glass protective scale. The most recent frontiers in a research oriented to high Technology Readiness Level (TRL) applications of the UHTC technology to aerospace involve the enhancement of mechanical properties by introducing short and continuous carbon fibre reinforcements in a UHTC matrix, leading to the definition of the Ultra-High-Temperature Ceramic Matrix Composites (UHTCMCs). The overall objective is developing large ultra-refractory aerospace transportation systems components with outstanding ablation resistance and enhanced mechanical properties and reliability. To achieve this goal, testing in a relevant environment is required to properly characterize the ceramic materials in conditions representative of the real flight applications. In this framework, the European Horizon 2020 C3HARME Project (Next Generation Ceramic Composites for Combustion Harsh Environments and Space), involving several research institutions and private companies all over Europe, has been focused on the development of a new class of UHTCMC materials, based on an UHTC matrix and reinforcing carbon fibres. The aim was to develop near-zero ablation Thermal Protection Systems and near-zero erosion hybrid and solid rocket nozzles. The University of Naples "Federico II" has been involved in the experimental campaigns aimed at the characterization of the new materials in relevant environments, in the definition of the geometries of the samples to be tested, and in numerical modelling of the aero-thermo-chemical conditions around the test articles and the material thermal response. Most of the activities described in the present thesis have been carried out in the framework of the C3HARME project. Specifically, the objectives of this work include: i) a characterization, as comprehensive as possible, of the aero-thermo-dynamic behaviour of ceramic materials, with different formulations, sizes and shapes, in lab-reproduced representative environments, and ii) the definition and employment of numerical models for the simulation of the thermo-fluid-dynamic conditions and material thermal response. After a general introduction to the topic and the description of experimental facilities and computational models, results obtained in the framework of materials characterization for hypersonic TPS are presented. Relevant tests were performed in an arc-jet supersonic wind tunnel, where the typical aero-thermodynamic and chemical conditions of atmospheric re-entry are reproduced at supersonic Mach numbers, temperatures above 5000 K and a consistent amount of dissociated oxygen and nitrogen. Non-intrusive diagnostic techniques (two-colour pyrometers and infrared thermo-cameras) were used to continuously monitor samples surface temperatures during testing. The main activities included: an extensive experimental campaign for the characterization of UHTCMC materials with almost zero ablation properties at ultra-high temperature; a series of dedicated test campaigns carried out in order to understand the effect of SiC content on the ultra-high-temperature oxidation and material behaviour under simulated re-entry conditions; specific experiments to compare the material response in different chemical atmospheres (simulated air and pure nitrogen flows); large scale tests on UHTCMC samples carried out in the arc-jet wind tunnel available at the premises of the German Aerospace Centre (DLR) in Cologne (Germany). The outcomes of all these experimental activities are presented and discussed, also in light of the post-test characterizations carried out, in collaboration with project partners, to investigate the features of the materials microstructures after the exposure to the atmospheric re-entry environment. Moreover, the experimental results were complemented by Computational Fluid Dynamics (CFD) simulations, employed to allow accurate prediction not only of the thermo-fluid-dynamic flow field around the test articles, but also of the thermal behaviour of the materials samples, including an investigation of the effect of material properties, such as thermal conductivity, emissivity and catalycity. A great interest has been paid to the interpretation of a phenomenon observed in several tests, consisting in a rapid temperature increment at constant flow conditions (known as temperature jump in the relevant literature regarding UHTCs and SiC-based ceramics). This thesis intends to propose a thorough and detailed analysis of the materials aero-thermo-dynamic behaviour at ultra-high temperatures in a representative re-entry environment, aiming to provide a comprehensive interpretation of the temperature jump, correlating the outcomes of infrared temperature measurements, post-test microstructural analyses and numerical simulations to highlight the parameters which mainly affect the heat transfer from the flow to the ceramic. The second part of results is focused on characterization of the same class of materials, for near-zero erosion rocket nozzles. The experimental activities were carried out with a 200 N-class hybrid rocket engine in different test configurations. The first tests have been performed with a novel, dedicated test set-up exposing UHTCMC samples to the supersonic exhaust jet of the hybrid rocket operated with gaseous oxygen burning cylindrical port High-Density PolyEthylene (HDPE) grains. Also in this case, non-intrusive diagnostic equipment has been employed to monitor the surface temperature of the samples. The combination of combustion temperature over 3000 K, supersonic Mach number and stagnation pressures allowed reproducing realistic rocket nozzles operating conditions, in order to demonstrate the ability of the specimens to preserve their functional integrity in a relevant environment. The second set-up ("Chamber insert test") was meant to assess the capability of the test article (in the shape and size of an annular combustion chamber element) to withstand high thermo-mechanical stress at high pressures in relevant aero-thermo-chemical combustion environment. As a third step, UHTCMC nozzle throat inserts have 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. Finally, complete subscale nozzles, made by different processing routes, matrix formulations, fibres architectures and having different properties (such as fibre volume content and porosity), were repeatedly tested in order to select the most promising technology for reusable non-eroding rocket nozzles. Also in this case, CFD models were exploited to characterise the flow field in the different test configurations, evaluate the aero-thermo-dynamic loads on the prototypes during tests and to rebuild the thermal response of the materials. Combining experimental and numerical results, general conclusions were drawn about the effect of different parameters, such as matrix composition, fibre architecture and volume content, densification process and final material porosity