8 th European Symposium on Aerothermodynamics for space vehicles HYBRID ROCKET TECHNOLOGY IN THE FRAME OF THE ITALIAN HYPROB PROGRAM M. Di Clemente, R. Votta, G. Ranuzzi, F. Ferrigno March 4, 2015
Outline 1. Introduction HYBRID Project overview 2. Design of the hybrid rocket engine Injection system Combustion chamber Nozzle 3. Numerical tool development 4. Conclusions Use or disclosure of the information contained herein is subject to specific written CIRA approval 2
HYPROB HYPROB program launched in 2011 under contract by the Italian Ministry of Education and Research has the objective to consolidate national capabilities on the overall engine system for future space applications. For both liquid LOx-CH 4 and hybrid paraffin-based developments, the program is structured in three main lines: Systems: design and development of technology demonstrators, including intermediate breadboards; Technology: R&D development in the areas of CFD combustion modeling, thermo-mechanical modeling and materials, advanced optical diagnostics; Experimental: testing capabilities for both basic physics and systemoriented (demonstrators) experimentation. The definition of the program, carried out by CIRA, involved institutional, industrial and scientific national stakeholders. Use or disclosure of the information contained herein is subject to specific written CIRA approval 3
HYBRID Project Hybrid is one of the implementing projects of HYPROB program The main objectives of the project are: Design and testing of a hybrid rocket engine as technology demonstrator of research activities, based on paraffin and nitrous oxide (N 2 O), thrust 30kN, re-ignition and throttability capabilities Research and development on alternative oxidizers as O 2 and H 2 O 2 Technology development of solid grain manufacturing process Development of a numerical platform to increase design tools capabilities able to simulate flow phenomena within a hybrid combustion chamber Use or disclosure of the information contained herein is subject to specific written CIRA approval 4
HRE Demonstrator Use or disclosure of the information contained herein is subject to specific written CIRA approval 5
Demonstrator dimensioning; Trade-off analyses More than 600 configurations have been analyzed by varying chamber pressure, grain length and port diameter The final configuration has been chosen to maximize Isp DESIGN DRIVERS Thrust : 30 kn Firing time 30 s + 5s as margin SELECTED CONFIGURATION Length = 0.8 m Internal Diam. = 0.1 m External Diam. = 0.35 m P_chamber = 40 bar Use or disclosure of the information contained herein is subject to specific written CIRA approval 6
Performances Estimation of performances based on 0-D approach assuming an ideal combustion A slight decrease of thrust and chamber pressure is predicted during the burning time (about 5%) Isp is almost constant because for the selected grain length the effect of internal diameter is negligible Present results do not consider throat regression due to ablation Use or disclosure of the information contained herein is subject to specific written CIRA approval 7
Injection System A fixed injector geometry + Fluid Control Valve has been selected through the Analytical Hierarchy Process Preliminary design has been verified through numerical computations in cavitating conditions Predicted 1-D mass flow rate: 222.5 g/s Selected injector diameter: 4 mm Liquid phase Gaseous phase CFD Results Injection Pressure (MPa) Number of injectors: 50 Chamber Pressure (MPa) Numerical Mass Flow Rate (g/s) Mod. Bernoulli % Error 5.5 4.0 212.4 < 5 Highlights Good overlap between analytical formula and CFD Cavitation starts at the injector inlet Steep phase change at injector exit Use or disclosure of the information contained herein is subject to specific written CIRA approval 8
The influence of the following parameters has been investigated by means of CFD: convergent length convergent angle radius of curvature at throat divergent angle Mach number Nozzle (1/2) Nozzle ID-25 Theta div: 18 deg Rc: 1.5 Rt Theta conv: 45 deg Selected configuration is a compromise between two opposite effects: Higher throat heat fluxes with lower total thermal loads lower throat heat fluxes with higher total thermal loads Use or disclosure of the information contained herein is subject to specific written CIRA approval 9
Throat Heat Flux: ~ 30 MW/m 2 Nozzle Cooling Problem Nozzle (2/2) Ablative graphite Nozzle A simplified approach based on CMA software with proper Scale factor has been used to evaluate the regression rate and the conductive heat flux t = 30s : T max =2765.57 K Thermal computation of the graphite nozzle Use or disclosure of the information contained herein is subject to specific written CIRA approval 10
Design assessment The overall design has been assessed through a complete CFD simulation of the motor configuration Domain Three different geometries corresponding to different instants during the burning time: 1. no grain recession 2. 7 seconds grain recession 3. 15 seconds grain recession Modeling Transient 2 order Ideal gas k-ω SST Turbulence Model EDM mono-step 35 bar 23 bar p CC = 36.6 bar results for the initial configuration without grain recession Flame does not occur at injector exit nor in the pre-chamber Estimated chamber pressure close to the design value (Pcc = 40 bar) Use or disclosure of the information contained herein is subject to specific written CIRA approval 11
Numerical Code Development Use or disclosure of the information contained herein is subject to specific written CIRA approval 12
Numerical demonstrator Software platform shall able to simulate the flow field features within the combustion chamber of an hybrid engine considering different oxidizers Simulation capability set of physical and numerical models suited for the Hybrid Propulsion CFD simulation Performance software platform oriented to the CFD support to the design activities expected software performances in terms of accuracy and computational costs License Policy CIRA CFD software (developed in-house) open-source CFD software 13
Software simulation capability The physical and numerical models required for the simulation of the internal flows in a hybrid rocket engine have been identified, as result of a literature review: Steady/Unsteady RANS formulation Up to 2 order time integration schemes Up to 2 order convective schemes Generic gas mixtures input interface Real gas models for the simulation of high pressure effects Turbulent combustion models Ignition models Two equation turbulence models Regression rate models for hybrid rocket applications 14
Software performances Three different levels of complexity have been identified for the CFD performance evaluation : Level-1: CFD test to support the HRE pre-design activity; Level-2: CFD test to support the HRE design activity or the numerical rebuilding of experimental tests; Level-3: CFD test to support the HRE design activity or the numerical rebuilding of experimental tests in the presence of complex geometries;
Software Platform CIRA-HYBRID The selected platform includes two different CFD solvers OpenFOAM NExT Open Field Operation and Manipulation Open-source CFD solver The Numerical Experiment Tool CIRA CFD solver NExT OpenFOAM CFD-HYBRID The software platform will integrate the Source code availability Yes Yes Yes simulation capabilities of the two Programming Yes Stiff Yes selected solvers Parallel computing Yes Yes Yes High Mach flow Yes Stiff Yes Low Mach flow Stiff Yes Yes Unstructured Grid No Yes Yes Multiphase flow No Yes Yes
Validation test cases Methane/Air turbulent diffusion flame [Ref. Brookes & Moss] Good comparison between numerical and experimental data at different locations Difference between numerical and experimental results is lower than 10% (on the maximum value) Use or disclosure of the information contained herein is subject to specific written CIRA approval 17
Validation test cases Hybrid rocket engine simulation [Grosses s hybrid rocket laboratory motor configuration] Pcc exp = 26.4 bar num = 24 bar Good comparison between numerical and experimental prediction of Pcc Good comparison between Fluent and Next results in terms of wall heat flux Use or disclosure of the information contained herein is subject to specific written CIRA approval 18
Conclusions HYBRID project is focused on the development of enabling technologies for hybrid rocket propulsion Developed technologies will be integrated into a ground demostrator, based on paraffin and nitrous oxide, whose preliminary design has been presented A numerical code able to predict flow phenomena within an hybrid combustion chamber is being developed The project is approaching the Preliminary Design Review and subsequently small and medium scale tests will be carried out to validate the design For further information: Marco Di Clemente m.diclemente@cira.it Use or disclosure of the information contained herein is subject to specific written CIRA approval 19