Experimental Firing Test Campaign and Nozzle Heat Transfer Reconstruction in a 200 N Hybrid Rocket Engine with Different Paraffin-Based Fuel Grain Lengths

Abstract

Firing test campaigns were carried out on a 200 N thrust-class hybrid rocket engine, using gaseous oxygen as an oxidizer and a paraffin-wax-based fuel. Different fuel grain lengths were adopted to extend the fuel characterization under different operating conditions, and to evaluate rocket performances and internal ballistics in the different configurations. In addition to data collected under a 220 mm propellant grain length, two further test campaigns were carried out considering 130 mm and 70 mm grain lengths. Two different injector types were adopted in the 130 mm configuration; in particular, a showerhead injection system was used with the aim to contain high-amplitude pressure oscillations observed during some firing tests in this engine configuration. Parameters such as the chamber pressure and temperature inside the graphite nozzle, space-averaged fuel regression rate and nozzle throat diameter were measured. The results allowed for the investigation of different issues related to hybrid rockets (e.g., fuel regression rate, engine performance, nozzle ablation under different conditions). The focus was mainly directed to the nozzle heat transfer, through the reconstruction of the convective heat transfer coefficient for different tests in the 70 mm grain length engine configuration. The reconstruction took advantage of the experimental data provided by the nozzle embedded thermocouple. Then, the experimental convective heat transfer coefficient was used to validate the results from some empirical correlations. The results showed significant differences between the experimental convective heat transfer coefficients when considering tests with different oxidizer mass flow rates. Furthermore, the predictions from the empirical correlations proved to be more reliable only in cases characterized by oxidizer-rich conditions.

Combustion of Date Stone and Jojoba Solid Waste in a Hybrid Rocket-like Combustion Chamber

“…Elmay et al [18] The expected propulsion performances of the fuels were estimated using the Chemical Equilibrium Analysis (CEA) code developed by NASA [28]. Considering the enthalpy of formation of −967.8 kJ/mol [29] for paraffin-based fuel, −378.72 kJ/mol (28.55 MJ/kg [14]) for date seed, and −95.89 kJ/mol (15.34 MJ/kg [30]) for jojoba, the estimated theoretical combustion temperature for the propellant combination of paraffin/oxygen, jojoba/oxygen, and date/oxygen under the frozen chemical equilibrium condition for a combustion pressure of 2 bar and 10 bar is shown in Figure 4. The maximum combustion temperature at 2 bar was 3384 K, 3406 K, and 3205 K for jojoba, date, and paraffin, respectively, indicating that the biomass fuels have a higher maximum combustion temperature.…”

Tests and simulations on 200N paraffin-oxygen hybrid rocket engines with different fuel grain lengths

“…Finally, measurement of temperature inside the graphite nozzle allowed for a rebuilding of wall heat flux in the S configuration (the most oxidizing) [4]. An iterative procedure was used to determine a profile of the convective heat transfer coefficient hc matching the experimental temperature measurement, by solving the unsteady energy equation inside the nozzle with CFD simulations.…”

“…In addition to data collected with 220 mm propellant grain length [ 3 ] (labeled as L), two further test campaigns were carried out considering 130 mm (labeled as M) and 70 mm (labeled as S) grain lengths. Full details on the experimental setup and results are reported in [4]. In this work, based on measurements of pressures, temperatures, thrust and mass flow rate, and with the support of Computational Fluid Dynamic (CFD) simulations, some considerations on performance, regression rate and graphite nozzle heat transfer.…”