1,720,965 research outputs found
Combined fluid-dynamic modelling of hybrid rocket internal ballistics and nozzle heat transfer
No full textA computational thermo-fluid-dynamic model of the hybrid rocket internal ballistics has been developed in the present work. Numerical simulations of the flowfield in a laboratory 200 N-class hybrid rocket engine, operated with gaseous oxygen and high-density polyethylene or acrylonitrile-butadiene-styrene 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 behaviours of graphite compared to ceramic material nozzle, highlighting the severe thermal gradient occurring in the ceramic material
Computational Evaluation of Aero-Thermo-Dynamic Loads and Effect of Catalyticity in an Arc-Jet Wind Tunnel
No full textWith 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, in order 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 as per 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. Due to 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, in agreement with what reported in literature
Computational fluid-dynamic modeling of the internal ballistics of paraffin-fueled hybrid rocket
No full textComputational fluid dynamics is becoming a key tool for reducing the hybrid rocket operation uncertainties and development cost, but numerous challenges, due to the complexity of modeling the solid fuel consumption mechanism and the interaction with the reacting flowfield, have still to be addressed. These latter features are further complicated with paraffins for the melted-fuel entrainment phenomenon. This paper presents a computational thermo-fluid-dynamic model of the internal ballistics of hybrid rockets burning gaseous oxygen and paraffin-based fuel. With the purpose of predicting the local fuel regression rate, the model is coupled with an improved gas/fuel-surface interface treatment based on local mass, energy and mean mixture fraction balances, combined to an additional analytical equation for the calculation of the entrainment fraction of the fuel consumption rate. Parametric analyses are carried out to assess the effect of fuel properties on the regression rate. Several experimental test cases, obtained from static firing of a laboratory-scale rocket, are simulated. Calculated regression rates show an error with respect to the measured data around 10% in the worst case. Chamber pressure is predicted with lower accuracy, with errors less than 20%; the main factor of deviation is shown to be the estimation of the combustion efficiency
A simplified computational fluid-dynamic approach to the oxidizer injector design in hybrid rockets
No full textFuel regression rate in hybrid rockets is non-negligibly affected by the oxidizer injection pattern. In this paper a simplified computational approach developed in an attempt to optimize the oxidizer injector design is discussed. Numerical simulations of the thermo-fluid-dynamic field in a hybrid rocket are carried out, with a commercial solver, to investigate into several injection configurations with the aim of increasing the fuel regression rate and minimizing the consumption unevenness, but still favoring the establishment of flow recirculation at the motor head end, which is generated with an axial nozzle injector and has been demonstrated to promote combustion stability, and both larger efficiency and regression rate. All the computations have been performed on the configuration of a lab-scale hybrid rocket motor available at the propulsion laboratory of the University of Naples with typical operating conditions.
After a preliminary comparison between the two baseline limiting cases of an axial subsonic nozzle injector and a uniform injection through the prechamber, a parametric analysis has been carried out by varying the oxidizer jet flow divergence angle, as well as the grain port diameter and the oxidizer mass flux to study the effect of the flow divergence on heat transfer distribution over the fuel surface. Some experimental firing test data are presented, and, under the hypothesis that fuel regression rate and surface heat flux are proportional, the measured fuel consumption axial profiles are compared with the predicted surface heat flux showing fairly good agreement, which allowed validating the employed design approach. Finally an optimized injector design is proposed
Transient Computational Thermofluid-Dynamic Simulation of Hybrid Rocket Internal Ballistics
No full textA computational thermofluid-dynamic model of hybrid rocket internal ballistics is developed. Numerical simulations of the flowfield in a laboratory small-scale hybrid rocket motor, operated with gaseous oxygen and high-density polyethylene propellants, are carried out with the aim of predicting the solid fuel regression rate experimentally achieved with two different oxidizer injectors. The fuel regression rate is the main parameter for the hybrid rocket design. Here, it is calculated with a detailed gas/surface interface characterization based on local mass and energy balances. The combustion of oxygen and gaseous ethylene injected from the fuel wall is modeled by means of the probability-density-function approach coupled to chemical equilibrium. Two oxidizer-injection configurations, which generate either a two-dimensional axially symmetric or three-dimensional flowfield, are analyzed. The local regression rate is evaluated along both the fuel grain axis and inner circumference in the three-dimensional case, as well as at several stages in the firing, by updating the local port diameter, which is thus not assumed constant. Data retrieved from three firing tests are compared with the numerical results, revealing good agreement between both the average regression rates (maximum deviation less than 5%) and the fuel consumption axial profiles (with maximum deviation of 14%
Computational Fluid-dynamic Simulations of Hybrid Rocket Internal Flow Including Discharge Nozzle
No full textA 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
Design, fabrication and high velocity oxy-fuel torch tests of a Cf-ZrB2- fiber nozzle to evaluate its potential in rocket motors
No full textThe resistance to ablation of a ceramic nozzle made of a 50 vol% Carbon fiber-50 vol% ZrB2 composite was evaluated in a high velocity oxy-fuel torch (HVOF) in conditions simulating typical exhaust engine flows. The composite was prepared by hot pressing a mixture of ZrB2 powders and C chopped fibers and characterized in terms of microstructural features, strength and toughness. Then, the sintered pellet was machined by electro-discharge machining to obtain a ceramic throat. The throat was assembled to convergent and divergent graphite parts to compose a segmented nozzle and tested in HVOF torch under a heat flux of 2.5 MW/m2 and flame temperature of 2730 K. Fluid dynamic simulations enabled to rebuild the heat field temperatures of the jet flow and of the solid nozzle. The throat well survived 30 s undergoing little oxidation of the frontal zone without dimension or shape variation. No appreciable ablation of the throat was measured
Effect of hypersonic flow chemical composition on the oxidation behavior of a super-strong UHTC
No full textA super-strong ZrB2 ceramic containing WC and SiC was tested in a supersonic arc-jet wind tunnel by exposure to flows with two chemical compositions, simulated air or pure nitrogen, at temperatures of 2650 and 2800 K. Temperature jumps of 500–600 K were observed in both environments at constant flow conditions. SEM analyses revealed that oxygen in the high-enthalpy flow retards the material consumption owing to the formation of partially protective glass. Then, it appears that even 5 vol% of W-compounds is sufficient to modify the subscale oxide configuration and form a cell-like multi-layered architecture typical of tungsten, rather than that typical of ZrB2
Design and Testing of a Paraffin-Based 1000 N HRE Breadboard
No full textThe 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
Experiments and simulations of hybrid rocket internal flows and material behaviour
Get PDFIn the last decade a significant and ever growing interest has been addressed towards hybrid rocket propulsion, which offers the best-of-both-worlds by leveraging the favourable aspect of both traditional solid and liquid systems. Among the numerous advantages which characterize hybrid rockets, the most attractive ones are the re-ignition and throttling capabilities combined with the possibility of embedding environmentally sustainable propellants and, of the utmost importance, their intrinsic safety and lower operational costs. Moreover, hybrid rockets yield a better specific impulse than solid propellant rockets and a higher density impulse than liquids, which make them a promising technology in a number of space missions.
The widely recognized potentialities of the hybrid rocket warrant the renewed research efforts that are being devoted to its development, but the state-of-the-art of this technology still presents a number of challenging issues to be solved.
A first fundamental task is the definition of suitable models for the prediction of the motor internal ballistics and performance. In particular, rocket performance is governed by the rate at which the fuel is gasified, i.e. by the fuel regression rate, as this latter determines the total mass flow rate and the overall oxidizer-to-fuel mixture ratio, which, for a given chamber pressure, control the motor thrust and the ideal specific impulse. For a given fuel, regression rate is basically limited by the heat flux input to the solid grain, which mainly depends on the thermo-fluid-dynamics in the combustion chamber. This latter is significantly influenced by several geometrical parameters, such as, for example, the oxidizer injection configuration or the grain port shape. Furthermore, the recent efforts aimed at overcoming the main drawback of the hybrid rockets, which is the low regression rate of conventional polymeric fuels, have been focused on the development of new paraffin-based fuels, characterized by a consumption mechanism presenting additional complex phenomena compared to that of conventional polymers. Their intrinsic characteristic is the onset of a thin liquid layer on the fuel grain surface, which may become unstable, leading to the lift-off and entrainment of fuel liquid droplets into the main gas stream, increasing the fuel mass transfer rate. This phenomenon is strongly susceptible to the fuel composition, its manufacturing process and the obtained thermo-mechanical properties as well as to the engine operating conditions, which makes the prediction of the regression rate and combustion chamber internal ballistics even harder than in the case of a pure polymer. In this framework, computational fluid dynamics of hybrid rocket internal ballistics is becoming a key tool for reducing the engine operation uncertainties and development cost, but its application still presents numerous challenges due to the complexity of modelling the phenomena involved in the fuel consumption mechanism and the interaction with the reacting flowfield, for both the cases of classical polymeric and liquefying paraffin-based fuels. A research effort is therefore of major importance in order to cover the lacking aspects and obtain quantitatively accurate results.
Another challenge for the hybrid rocket technology development is the optimization of the design of thermal insulations. The inner surface of the exhaust nozzle, through which the flow is accelerated to supersonic conditions producing the required thrust, is the most critical in this sense, as it is subjected to the highest shear stress and heat fluxes in a chemically aggressive environment. These severe conditions usually lead to removal of surface material due to heterogeneous reactions between oxidizing species in the hot gas and the solid wall. Because of the material erosion, there is an enlargement of the nozzle throat section and a consequent decrease of rocket thrust, with detrimental effects over the motor operation. Thus, the requirement that dimensional stability of the nozzle throat should be maintained makes the selection of suitable rocket nozzle materials extremely hard. In recent years, Ultra-High-Temperature Ceramics (UHTC) and Ultra-High-Temperature Ceramic Matrix Composites are the subject of considerable interest as innovative materials for rocket application, but still need to be properly characterized. Experimental testing along with computational fluid dynamic (CFD) simulations are, thus, both needed to improve the design and the current performance prediction capabilities of such propulsion systems. In this framework, the University of Naples is involved in the European project C3HARME – Next Generation Ceramic Composites for Combustion Harsh Environment and Space, in collaboration with other research centres, universities and industries, which aims at the design, manufacturing and testing of new-class high-performance UHTCMC for near-zero erosion rocket nozzles.
In the present work, the above-mentioned challenges are dealt with taking a combined experimental/numerical approach to improve understanding of the interaction between the gaseous combusting flow typical of hybrid rocket engines and the surface of solid materials involved in their operation, with a special focus to the fuel grain present in the combustion chamber, with the aim of predicting its consumption mechanism, and the exhaust nozzle inner surface, with the aim of identifying and validating new-class UHTCMC materials with improved erosion and structural resistance to the severe conditions experienced in particular in the throat region.
In particular, the first main objective of the present work is the definition of proper computational thermo-fluid-dynamic models of the hybrid rocket internal ballistics, including a dedicated gas/surface interface treatment based on local mass, energy and mean mixture fraction balances as well as proper turbulence boundary conditions, which can properly model the physical fuel consumption mechanism in both the cases of polymeric and liquefying fuels. For the validation of the computational models, a number of experimental test cases, obtained from static firing of laboratory scale rockets, have been performed at the Aerospace Propulsion Laboratory of University of Naples “Federico II” and successively numerically reconstructed. The comparison between the numerical results and the corresponding experimental data allowed validating the adopted model and identifying possible future improvements.
Then, the research activities for the characterization of new-class UHTCMC materials are presented and discussed. This part of the work was mainly focused on an extensive experimental campaign for the characterization of new-class UHTCMC materials. In particular, first preliminary tests on small samples exposed to the supersonic exhaust jet of a 200N-class hybrid rocket operated with gaseous oxygen burning cylindrical port High-Density PolyEthylene (HDPE) fuel grains have been carried out for a fast characterization and a preliminary screening of the best candidates for the final applications. After that UHTCMC nozzle throat inserts has been manufactured and experimentally tested to verify the erosion resistance and evaluate the effects on the rocket performance by comparison with those obtained in similar operating conditions employing a graphite nozzle. The experimental activities are supported by simplified low-computational-cost numerical simulations, whose main objectives has been the prediction of the complex flow field in the hybrid rocket combustion chamber and the thermo-fluid dynamic conditions on the material. Future research activities will be then focused to the further development of the numerical models with the extension of the treatment for the gaseous flow/solid surface interaction in order to get a deeper insight on the new materials behaviour
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