The paper presents some relevant achievements in hybrid rocket propulsion carried out by the Italian Aerospace Research Centre. On the basis of the experimental results obtained on a 200 N thrust class engine, a 1000 N class breadboard, fed with gaseous oxygen coupled with a paraffin-based fuel grain, was designed and experimentally tested in different conditions. The breadboard exhibited a stable combustion in all the firing test conditions; the testing campaign allowed the acquisition of different experimental data, as pre and post-combustion chamber pressure, throat material temperature, pre-combustion chamber temperature. The new breadboard was characterized by higher measured regression rate values with respect to corresponding data obtained with the smaller scale one, highlighting that the oxidizer mass flux is not the only operating quantity affecting the fuel consumption behavior, which could be also influenced by scale parameters, such as the grain port diameter, and other operating conditions, such as the mixture ratio. Aerospace 2019, 6, 89 2 of 15Detailed studies have been presented by Karabeyoglu [4,5], where it is demonstrated that the fuel composition and its thermo-mechanical properties strongly affect the liquid layer instability and, therefore, the fuel regression rate.In this scenario, the present research was mainly focused on the investigation of the fuel characteristics, in order to ensure high performance in terms of regression rate but maintaining good mechanical properties. In a previous authors' work, results of an experimental test campaign, performed at subscale level on a 200 N breadboard, demonstrated very good performances and mechanical properties of the analyzed paraffin-wax formulation [6].Based on these results, the main objective of the present work is the scale-up of the fuel grain, adopting the same fuel formulation, and the design of a new breadboard, moving towards the 1000 N thrust class. A new experimental test campaign was carried out on the 1000 N breadboard, allowing for the investigation of the paraffin-based fuel blend behavior on a larger scale. In particular, with respect to subscale experiments, the oxidizer to fuel mixture ratio range was extended. Numerous data were acquired, including chamber pressure, thrust, temperature of the flow in the pre-chamber and temperature inside the graphite nozzle material. The latter parameter allows for the estimation of the convective heat transfer coefficient in the nozzle region, which is strictly linked to the graphite nozzle thermo-chemical erosion. This is an extremely important parameter to be evaluated, since the throat area enlargement directly affects the motor performances [7].The results of the experimental test campaign show that hybrid rocket engine can operate, with good efficiency and stability, in a wide range of operating conditions, confirming some of the advantages over both solid and liquid technologies often mentioned in the relevant literature [1,2,8,9]. Methodology and Design Design LogicThe procedure, ad...

The successful design of a liquid rocket engine is strictly linked to the development of efficient cooling systems, able to dissipate huge thermal loads coming from the combustion in the thrust chamber. Generally, cooling architectures are based on regenerative strategies, adopting fuels as coolants; and on cooling jackets, including several narrow axial channels allocated around the thrust chambers. Moreover, since cryogenic fuels are used, as in the case of oxygen/methane-based liquid rocket engines, the refrigerant is injected in liquid phase at supercritical pressure conditions and heated by the thermal load coming from the combustion chamber, which tends to experience transcritical conditions until behaving as a supercritical vapor before exiting the cooling jacket. The comprehension of fluid behavior inside the cooling jackets of liquid-oxygen/methane rocket engines as a function of different operative conditions represents not only a current topic but a critical issue for the development of future propulsion systems. Hence, the current manuscript discusses the results concerning the cooling jacket equipping the liquid-oxygen/liquid-methane demonstrator, designed and manufactured within the scope of HYPROB-NEW Italian Project. In particular, numerical results considering the nominal operating conditions and the influence of variables, such as the inlet temperature and pressure values of refrigerant as well as mass-flow rate, are shown to discuss the fluid transcritical behavior inside the cooling channels and give indications on the numerical methodologies, supporting the design of liquid-oxygen/liquid-methane rocket-engine cooling systems. Validation has been accomplished by means of experimental results obtained through a specific test article, provided with a cooling channel, characterized by dimensions representative of HYPROB DEMO-0A regenerative combustion chamber.

The HYPROB Program, developed by the Italian Aerospace Research Centre, has the aim of increasing the Italian system design and manufacturing capabilities on liquid oxygen-hydrocarbon rocket engines; the most important activity is represented by the development and testing of a ground engine demonstrator of three tons thrust based on methane as propellant. The demonstrator baseline concept is featured by 18 injectors and is regeneratively cooled by using liquid methane. The cooling system has a counter-flow architecture and is made by 96 axial channels; methane enters the channels in the nozzle region in supercritical liquid condition, is heated by the combustion gases along the cooling jacket and then is injected into the combustion chamber as a supercritical gas. The goal of the present paper is to describe the activities supporting the cooling jacket design, aiming at identifying the optimal configuration of the cooling channels. 3-D CFD analyses have been performed on different cooling channel arrangements, in terms of channel height and rib width. Moreover, simulations described the thermo-fluid dynamic behavior of methane by means of NIST real gas modeling and they were necessary to give the proper input to the thermo-structural analyses in order to verify the most critical sections of the cooling jacket.

Reliability of liquid rocket engines is strictly connected with the successful operation of cooling jackets, able to sustain the impressive operative conditions in terms of huge thermal and mechanical loads, generated in thrust chambers. Cryogenic fuels, like methane or hydrogen, are often used as coolants and they may behave as transcritical fluids flowing in the jackets: after injection in a liquid state, a phase pseudo-change occurs along the chamber because of the heat released by combustion gases and coolants exiting as a vapour. Thus, in the development of such subsystems, important issues are focused on numerical methodologies adopted to simulate the fluid thermal behaviour inside the jackets, design procedures as well as manufacturing and technological process topics. The present paper includes the numerical thermal analyses regarding the cooling jacket belonging to the liquid oxygen/liquid methane demonstrator, realized in the framework of the HYPROB (HYdrocarbon PROpulsion test Bench) program. Numerical results considering the nominal operating conditions of cooling jackets in the methane-fuelled mode and the water-fed one are included in the case of the application of electrodeposition process for manufacturing. A comparison with a similar cooling jacket, realized through the conventional brazing process, is addressed to underline the benefits of the application of electrodeposition technology.

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