Giuseppe Daniele Di Martino

Transpiration cooling is a very promising active method for thermal management of critical components for hypersonic flight. On the other side, its efficiency is affected by the characteristic of the cooling gas outflow from the porous material. This paper presents a study on the outflow characteristics of a sharp-edged, transpiration cooled fin as flown as part of the FinEx experiment of the HIFLIER1 sounding rocket. The permeable ceramic matrix composite material used for the transpiration cooled leading edge shows in general a outflow pattern dominated by single jets, distributed relatively homogeneous over the surface, disrupted by significant strong jets induced by small inhomogeneities. The general strongly anisotropic outflow characteristics were also investigated. The findings of the outflow characterization led to valuable insights for the design of future sharp-edged transpiration cooled structures.

Re-entry flights of reusable first or upper stages typically foresee phases in the hypersonic flight regime, characterized by severe aero-thermal loads which could become critical for the most exposed components, like the vehicle forebody or the fin leading edges. These require consequently dedicated thermal protection systems (TPS), whose design generally requires a multi-disciplinary approach. In this framework, a viable solution is the use of high-temperature resistant ceramic matrix composite (CMC) structures, but the implementation of such technology, especially for the manufacturing of complex components and its application in real flight conditions, still presents significant challenges. In this work, the design activities for the CMC-based TPS of the payload forebody of a hypersonic sounding rocket are presented, developed within the framework of the STORT project, whose mission includes in flight demonstration of multiple critical technologies required for sustained flight at Mach numbers above 8, corresponding to a significantly high integral thermal load.

With a view to designing tests in the arc wind tunnel small planetary entry simulator (SPES) at the Department of Industrial Engineering of the University of Naples “Federico II” on specimens of ultra-high-temperature ceramic (UHTC) materials, computer simulations have been carried out in order to get information about the test conditions, i.e., electrical power supplied to the torch, the nozzle area ratio (exit area/throat area), and the specimen geometry, to avoid expensive and time-consuming experimental characterization. Computations have been carried out both by a computation fluid-dynamics (CFD) and a direct simulation Monte Carlo (DSMC) code. The computations provided important quantitative information about flow field for velocity, temperature, gas composition, etc., and about heat flux and pressure on the specimen. The present computations allowed the authors to define the capabilities of SPES to perform tests on UHTC specimens, identifying a range of potential applications. Computed values of heat flux have been also compared with preliminary measurements performed by means of a copper slug calorimeter. Because of the evident effect of surface catalytic level on the experimental measurements, an experimental/computational procedure has been used to estimate the recombination efficiencies of oxygen and nitrogen on the calorimeter surface. The definition of this procedure is aimed at a future evaluation of surface catalyticity of UHTC specimens. At the conditions tested in the present paper, the procedure indicates that oxygen catalytic effect is higher than that of nitrogen, which is in agreement with what is reported in the literature.

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.

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.

This work deals with experimental characterisation of ultra-high-temperature ceramic materials for aerospace applications. Current research activities are focused on development and testing of new-class ultra-high-temperature ceramic matrix composites for near-zero ablation Thermal Protection Systems and near-zero erosion rocket nozzles. To investigate materials behaviour in atmospheric re-entry conditions, relevant tests are carried out on samples with different shapes with a supersonic arc-jet facility, with specific total enthalpies higher than 20 MJ kg⁻¹ and temperatures over 2000 K in a gas atmosphere with high concentration of atomic oxygen. The Aerospace Propulsion Laboratory allows investigating rocket components or subsystems manufactured in innovative materials, such as nozzles and nozzle inserts, but also small specimens exposed to the rocket exhaust plume, in highly relevant operating conditions. Computational models for numerical simulations of high-enthalpy flows, in both arc-jet and propulsion environments, and thermal analyses of the material samples are developed to support the experimental activities.

A computational thermo-fluid-dynamic model of the hybrid rocket internal ballistics has been developed in the present work. Numerical simulations of the flowfield in two laboratory scales hybrid rocket motors, one in the 200-N class and the other in the 1 kN-class, operated with gaseous oxygen and high-density polyethylene have been carried out. The objective is twofold: first 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 as well as proper turbulence boundary conditions, along with chamber pressure and combustion efficiency. Second, the detailed study of the discharge nozzle flow and heat transfer. For the validation of the model, data retrieved from two firing tests are compared with the numerical results revealing good agreement of the average regression rates, fuel consumption axial profiles, and of the chamber pressure and combustion efficiency. The output of the motor ballistic simulations are then used for a detailed numerical study of the flow through the nozzle and of the unsteady thermal field inside the nozzle solid block showing different behaviors of graphite compared to ceramic material nozzle, highlighting the severe thermal gradient occurring in the ceramic material