During May 2012, the CERN-CNGS neutrino beam has been operated for two weeks for a total of ∼ 1.8 × 10 17 p.o.t., with the proton beam made of bunches, few ns wide and separated by 100 ns. This beam structure allows a very accurate time of flight measurement of neutrinos from CERN to LNGS on an event-by-event basis.Both the ICARUS-T600 PMT-DAQ and the CERN-LNGS timing synchronization have been substantially improved for this campaign, taking advantage of additional independent GPS receivers, both at CERN and LNGS as well as of the deployment of the "White Rabbit" protocol both at CERN and LNGS.The ICARUS-T600 detector has collected 25 beam-associated events; the corresponding time of flight has been accurately evaluated, using all different time synchronization paths.The measured neutrino time of flight is compatible with the arrival of all events with speed equivalent to the one of light: the difference between the expected value based on the speed of light and the measured value is δt = tof c −tof ν = 0.10±0.67 stat. ±2.39 syst. ns. This result is in agreement with the value previously reported by the ICARUS Collaboration, δt = 0.3 ± 4.9 stat. ± 9.0 syst. ns, but with improved statistical and systematic accuracy.
Numerical simulations of the flowfield in a gaseous oxygen/hydroxyl-terminated polybutadiene hybrid rocket engine are carried out with a Reynolds-averaged Navier–Stokes solver including detailed gas/surface interaction modeling based on surface mass and energy balances. Fuel pyrolysis is modeled via finite-rate Arrhenius kinetics. A simplified two-step global reaction mechanism is considered for the gas-phase chemistry to model the combustion of 1,3-butadiene in oxygen. Results are compared with the firing test data obtained from a laboratory-scale hybrid rocket in which gaseous oxygen is fed into axisymmetric hydroxyl-terminated polybutadiene cylindrical grains through an axial conical subsonic nozzle. With the oxidizer fed by this injector, which generates nonuniform conditions at the entrance of the fuel port, the fuel regression rate is shown to increase several times with respect to the case of homogeneous injection of the oxidizer through all the grain port area, in agreement with the experimental findings. The spatial distribution of the regression rate is substantially different from the one expected in a turbulent developing flow through a pipe because of the presence of a strong recirculation zone at the motor head end. Numerical simulations, although not capturing the absolute regression-rate values, are fairly able to predict the main ballistic features: that is, the regression rate’s weaker dependence on the mass flux, the influence of port diameter, and the pressure evolution over time, as well as the general trends of different firing tests
The coupled hot-gas/wall/coolant environment that occurs in regeneratively cooled liquid rocket engines is studied by a computational procedure able to provide a quick and reliable prediction of thrust-chamber wall temperature and heat flux as well as coolant-flow characteristics, like pressure drop and temperature gain in the regenerative circuit. The coupled analysis is performed by means of a CFD solver of the Reynolds-Averaged Navier-Stokes equations for the hot-gas flow and by a simplified quasi-2D approach, which widely relies on semi-empirical relations, for the coolant flow and wall structure heat transfer in the cooling channels. Coupled computations of the Space Shuttle Main Engine Main Combustion Chamber are performed and compared with available literature data. Results show a reasonable agreement in terms of coolant pressure drop and temperature gain with nominal data, whereas the computed wall temperature peak is closer to hot-firing test data than to the nominal value
Cooling of liquid rocket thrust chamber walls to allowable solid material temperatures induces near-wall chemical reactions, which are known to have an important role on the heat transfer from the hot gas to the wall. In this study, the contribution of near-wall chemical reactions to heat flux is investigated and quantified by suitable numerical analyses. Numerical results are first compared to literature experimental data of wall heat flux in subscale calorimetric thrust chambers for both oxygen/methane and oxygen/hydrogen propellant combination. Then, a parametric analysis is carried out varying chamber pressure, wall temperature, and propellant combination. This study highlights that oxygen/methane combustion products are more subject to near-wall recombination phenomena. They provide an increase of wall heat flux between 20 and 30% with respect to the frozen flow model evaluation, whereas in the case of oxygen/hydrogen, the wall heat flux increase due to recombination reactions is between 7 and 14%
Film cooling as a thermal protection for the walls of liquid rocket engines is studied numerically for hydrogen and methane thrust chamber tests. The aim is to verify the capability of the Reynolds Average Navier-Stokes model to capture the basic characteristics of film-cooled thrust chambers, considering a simplified approach, named pseudo-injector approach, which does not model propellant injection and combustion. This assumption allows a great saving in computational time, in particular when considering 3D simulations. The present study takes its origin from the European Community In-Space Propulsion 1 (ISP-1) program where, among various projects, an experimental campaign has been designed to study the film cooling technique in an oxygen/methane thrust chamber and to provide a database for computational fluid dynamics validation. The results show that the present approach gives good results in terms of heat flux characterization, in particular when dealing with test cases of high chamber pressure.
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