Francesco Battista

Design and Development of a LOX/LCH4 Technology Demonstrator

The HYPROB program is carried out by CIRA under contract by the Italian Ministry of Research with the main objective to improve National system and technology capabilities on liquid rocket engines (LRE) for future space applications, with specific regard to LOx/LCH 4 technology. Among its objectives are the design and development of technology LRE demonstrators, including intermediate breadboards. The current design of the demonstrator project along as the development plan, shall be hereinafter presented and discussed.

Aerothermal Environment Definition for a Reusable Experimental Re-entry Vehicle Wing

The development of the wing leading edge thermal protection system for a re-entry vehicle requires a deep understanding of the aero-thermal environment surrounding the vehicle. A technological project, founded by the Italian Space Agency, focused on the assessment of the applicability of Ultra High Temperature Ceramics (UHTC) or active cooling system to the fabrication of high performance and slender shaped hot structures for reusable launch vehicles wing leading edge has been started considering the configuration and the mission profile of the future experimental Flight Test Bed, named FTB-X, whose analysis is going on in the framework of the USV Program carried out at CIRA, the Italian Aerospace Research Centre within the P.R.O.R.A (Italian Aerospace Research Program). The current activity has been carried out within the Advanced Structural Assembly (ASA) project funded by the Italian Space Agency and by Thales Alenia Space as prime. Aim of this paper is to show the development and application of a simplified methodology to evaluate the time history of the aerothermal environment over the wing, capable of giving results as much reliable as possible in a reasonable time. This methodology is mainly based on the combined use of bi-dimensional CFD simulations over a number of wing sections suitably scaled by means of a restricted set of three dimensional CFD simulations. Results of the methodology has been successfully compared with full vehicle computations in thermochemical non equilibrium.

An Extrapolation from Flight Methodology for a Re-Entry Vehicle Wing Leading Edge Test in a Plasma Wind Tunnel Facility

Within the frame of the Advanced Structural Assembly (ASA) project funded by the Italian Space Agency and with Thales Alenia Space as prime, CIRA (Italian Aerospace Research Centre) will perform a series of Plasma Wind Tunnel Test in its facility “Scirocco” over a Test Article representative of the wing of the future experimental Flight Test Bed, named FTB-X. The validation and the qualification of thermal protection system technologies, developed to face repeated re-entry missions or a sustained hypersonic velocity of a space vehicle, can be done in a plasma wind tunnel by reproducing the flight thermal loads on a representative model. The on-ground high enthalpy facilities do not allow the simultaneous reproduction of all the thermo-fluid-dynamics conditions that characterize above all the low-earth orbit part of a typical space vehicle re-entry trajectory because it may be difficult to contemporary reproduce both heat flux and pressure. Moreover, the correct reproduction of heat flux and pressure at the stagnation point do not assure in general that we are simulating the same environment downstream of the stagnation point itself with respect to the flight conditions because of differences of the unit Reynolds number and Mach number. For all the above reasons an extensive theoretical-numerical analysis is necessary both for the extrapolation from simulated flight conditions to suitable plasma wind tunnel operating conditions (extrapolation-from-flight ), and for the extrapolation of the test results to flight conditions (extrapolation-to-flight ). The developed extrapolation from flight methodology, presented in this paper, has been applied to the design of a test campaign on a large scale wing leading edge model. Aim of this test campaign is to qualify different innovative TPS concepts with an aero-thermal environment as much as possible representative of the flight conditions that vehicle FTB-X will experience during the re-entry mission.

Cooling channel analysis of a LOX/LCH4 rocket engine demonstrator

A computational procedure able to describe the coupled hot-gas/wall/coolant environment that occurs in most liquid rocket engines is presented and demonstrated. The coupled analysis is performed by loose coupling of the two-dimensional axisymmetric ReynoldsAveraged Navier-Stokes equations for the hot-gas flow and the conjugate three-dimensional model for the coolant flow and solid material heat transfer in the regenerative cooling circuit. The latter model is in turn based on the coupled Reynolds-Averaged Navier-Stokes equations for the coolant flow and Fourier equation for the thermal conduction in the solid material. In this study, the thermal behavior of a regeneratively cooled oxygen/methane engine demonstrator is analyzed in detail. Starting from a nominal operative condition of the engine, different levels of channel surface roughness and coolant mass flow rate are considered in order to understand their influence on the heat transfer capability of the cooling system. Results show that the heat transfer can be markedly impaired if the operating parameters undergo rather minor changes with respect to the nominal condition.

Analysis of a low-emission combustion strategy for a high performance trans-atmospheric aircraft engine

This work describes the CIRA activities within the EU LAPCAT II project concerning the design of an air-breathing, hydrogen-fed, low-NO x combustion chamber for hypersonic cruise at Mach 5. The combustion, at the design equivalence ratio ER=0.8 in a conventional subsonic combustor, leads to an unacceptably high NO x generation (about 5-6000ppm), which has a strong impact on the ozone layer depletion at high altitude. Therefore, an advanced combustion strategy, namely the Rich-burn Quick-mix Lean-burn (RQL) combustion system, has been selected in order to achieve a reduction in NO x emissions. This strategy allows changing the local ER, moving combustion stages far from the stoichiometric ER but still preserving the overall ER and, consequently, the engine performance. The RQL combustion strategy adopts a two-staged combustion: in the first stage, all the fuel is injected and reacts with a fraction of the total air-flow, thus resulting in a fuel-rich combustion; in the second stage, the remaining air is mixed with the main flow and reacts with the remaining fuel in a fuel-lean combustion.

Supersonic Combustion Models Application in Advanced Propulsion Concepts

§currently at ALSTOM, Brown Boveri Strasse 7, Baden CH-5401, Switzerland This paper presents some applications of the CIRA 3D code C3NS-DB in the simulation of typical problems of supersonic combustion. After a brief code description, a trade-off analysis, carried out in order to find a reduced kinetic mechanism for Air-H2 combustion, is presented. The selected mechanism has to be computationally not heavy but, at the same time, accurate in prediction of ignition delay time and adiabatic flame temperature. Furthermore, 2D and 3D flow problems reproducing supersonic mixing and combustion processes have been investigated with the purpose of obtaining a verification of the implemented models and a validation by comparing numerical and available experimental data. In particular, four tests, of growing complexity, representative of typical supersonic combustion configuration have been selected. The first two ones are typical validation tests simulating respectively the mixing process in a normal-injection configuration (TC-1) and both mixing and combustion in a parallel-injection configuration (TC-2). The second two tests are typical scramjet engine applications: the TC-3 is a test case representative of a scramjet combustor and TC-4 is a scramjet wind tunnel model in fuel-off conditions. The results show a good agreement with experimental data and are encouraging in order to employ the code in more complex applications including the evaluation of propulsion systems performances.

Validation of Conjugate Heat Transfer Model for Rocket Cooling with Supercritical Methane

A numerical solver able to describe a rocket engine cooling channel fed with supercritical methane is validated against experimental data coming from a test article conceived and tested by the Italian Aerospace Research Center. The multidimensional conjugate heat transfer model numerically solves the Reynolds-averaged Navier–Stokes equations for the coolant flow and the Fourier’s law of conduction for the heat transfer within the wall. In this study, an experimental test case is reproduced in detail in order to evaluate the influence of partially unknown parameters, such as surface roughness and wall thermal conductivity, and of operative parameter uncertainty, such as the coolant mass flow rate and input heat transfer rate. The comparison made with respect to the wall temperature and coolant pressure drop of the whole set of experimental data provides complementary information that allows better understanding of experiments and infers possible deviations from the expected behavior.

Design of an Experimental Campaign on Methane Regenerative Liquid Rocket Engine Cooling System

The use of the Methane as coolant in a regenerative liquid rocket engine (LRE) presents some difficulties since transcritical fluidynamics operating conditions occur in the cooling channels. Transcritical conditions cause large fluid properties variation that strongly influences the coolant performance. The HYPROB program is carried out by CIRA under contract by the Italian Ministry of Research with the main objective to improve National system and technology capabilities on liquid rocket engines for future space applications, with specific regard to LOx/LCH4 technology. Among its objectives are the design and development of technology LRE demonstrators, including a specific breadboard based on an electrical heating to validate numerical methodologies and to improve the understanding of relevant physics of methane thermal properties in transcritical conditions.

Experimental investigations on the magneto-hydro-dynamic interaction around a blunt body in a hypersonic unseeded air flow

This paper deals with the experimental investigation on the MHD (magneto-hydro-dynamic or magneto-fluid-dynamic) interaction around a test body immersed into a hypersonic unseeded air flow. The experiments have been carried out in the CIRA plasma wind tunnel SCIROCCO. Two test conditions have been utilized for the experiments with a total pressure of 2.5 and 2.3 bar respectively, a total specific enthalpy of 16 and 12.1 MJ/kg respectively. The air flow was accelerated in the nozzle up to Mach 10. The magnetic induction field is generated by an electromagnet enclosed in the test body and reaches a 0.8 T maximum value in the interaction region.

Numerical Rebuilding of MHD Tests In An Unseeded Mach 10 Air Flow around a Blunt Body

These paper deals with the numerical rebuilding of the experimental data from the experimental campaing on blunt body equipped with an electromagnet made in CIRA at SCIROCCO facility. A set resticted set of experimental condition have been considered for numerical rebuilding. Two different CIRA codes have been used, CAST and EMC3NS. Codes, whose structure is generally similar, are different is some modeling aspects. Their results compared with experimental data have been analized and criticallycompared, and gives a guidelines for further developments and analysis in this field.