Injector Plate Research Articles

Two-Stage Impinging-Jet Injector Flow Dynamics and Mixing: Kerosene and Hydrogen Peroxide Propellants

ABSTRACT Kerosene/hydrogen-peroxide (H2O2) green propellant system has relatively high operational ratio of oxidizer (O) to fuel (F) that complicates the injector design and its layout. In this research, a procedure of two-stage, like-doublet (O-F-F-O) impinging-jet injector design is presented, where mixing of H2O2 and kerosene occurs between O-O and F-F sprays. Investigations at O/F ratios between 3.75 and 6.25 are performed. The impinging distance, angle, and jet velocity of the two-stage impinging-jet module are the design parameters examined. The PLIF technique is used to observe the spatial droplet distribution of the individual spray. The predicted mixing behavior by overlapping the individual fuel and oxidizer sprays is compared to the actual mixing of the O-F-F-O spray to justify the spray/spray interactions. Detailed behaviors of the fuel and oxidizer sprays are revealed. The effects of O/F ratio and other design parameters on the injector flow dynamics and propellant mixing are examined systematically. With experimentally defined shape factors, an optimum design of the injector element can be determined for a given injector plate. The design of the injector plate for a 450 N thrust engine is demonstrated, and the design is justified by the hot-fire test of the engine.

Investigation of full-scale porous injector plate transpiration cooling coupled with combustion in high-thrust H2/O2 rocket engines

The high chamber pressure and large heat flux impose greater demands on the structure cooling and thermal protection of the future high-thrust H2/O2 rocket engine. To investigate the applicability of transpiration cooling on the full-scale injector plate in the high-thrust rocket engine, a numerical model is presented, in which the porous region and the mainstream hot gas region are directly coupled. The local equilibrium model and eddy dissipation model are considered. The numerical model is validated by transpiration cooling experiments for a subscale rocket engine injector plate. Then, the numerical simulation of transpiration cooling for the high-thrust H2/O2 rocket engine injector plate shows that the transpiration cooling with 13% of the total hydrogen mass flow rate can reduce the gas side temperature of the injector plate by about 150 K and reduce the thermal penetration depth to 9.2% of the plate thickness. Moreover, As the coolant mass flow rate increase from 4% to 13%, the gas side temperature of the plate decreases from 735 K to 668 K and the cooling effectiveness rises. The lower the coolant inlet temperature, the lower the plate temperature, but the higher the cooling effectiveness. Besides, the plate temperature drops with the decreases of the thermal conductivity of the porous media.

Design and Analysis of Injector with Different Configurations for Liquid Rocket Engine

The performance of the rocket is determined by atomization, mixing and combustion process. Injectors are used to control these processes. Suitable method should be followed, so that the performance should be increased. The injector implementation in rockets determines the performance of the nozzle that can be achieved. Injectors can be as simple as a number of small diameter holes arranged in carefully constructed patterns through which the fuel and oxidizer travel. The performance of an injector can be improved by either using a superior propellant combustion, increasing the mass flow rate, positioning the angle of the hole or by reducing the size & increasing the number of orifices on the injector plate. In here the position is varied for different angles for injectors to improve the performance of the rocket. The injector is modeled using the Solidworks and a two dimensional combustion analysis to be carried out using ANSYS Fluent for different cases.

Effects of Flow Parameters on the Characteristics of Biogas Flames in a Crossflow Air Stream

This study presents the flame characteristics of laminar cross-flow non-premixed biogas-air flames established in a boundary layer using ANSYS Fluent. The validated numerical model is comprehensive enough to include variable thermo-physical properties, multi-component diffusion, short chemical kinetics mechanism, and optically thin approximation-based radiation model. Fuel is injected at a uniform velocity and the air is supplied perpendicular to it. Predicted and the in-house experimental biogas flames are qualitatively compared in terms of flame extents. The fuel flow rate has been varied from 0.02 m/s to 0.04 m/s to study the effects of power rating on the flame characteristics. At a given fuel flow rate, the airflow rate is incremented in steps to study the effects of forced convection on the characteristics and stability of the flames. Furthermore, the effects of the length of the flat plate upstream of the leading edge of the fuel injector plate on the flame anchoring location have been studied. The structure and characteristics of the reactive boundary layer are explained by using fields of temperature, velocity, mass fractions of major species, stoichiometric equivalence ratio, and Damköhler number. The reactive boundary layer fueled by methane is compared with that of biogas.

Numerical analysis of the hybrid rocket motor with axial injection based on oxidizer flow distribution

This paper studies the effects of liquid oxidizer flow distribution on combustion characteristics and performance of hybrid rocket motor with axial injection through numerical simulations of two-phase flow. Axial injection is widely used in hybrid rocket motor, however, the regression of burning surface may change combustion efficiency of the motor at different moments. Consequently, it is necessary for hybrid rocket motor to design injector patterns that could increase the average combustion efficiency in the whole operating process. The orifice injector plate is divided into three areas, which are recorded as zone I, zone II and zone III, respectively. Oxidizer flow distribution is achieved through changing the diameter of injector holes in various zones, and three-dimensional steady numerical simulations of two-phase flow of hybrid rocket motor with axial injector patterns at different moments are conducted. The motor adopts 98% hydrogen peroxide and hydroxyl-terminated polybutadiene as the propellants. Numerical analysis reveals that oxidizer flow distribution mainly influences the flow field structure in the pre-combustion chamber and front part of fuel grain port. Increasing oxidizer mass flow rate in zone III is beneficial for the chamber head thermal environment. Relative position in the radial direction of fuel grain inner surface and oxidizer distributed in the injector plate affects fuel regression rates, and the influence mainly exists in the front part of solid fuel grain. When the oxidizer mass flow rate in zone I is equal, the average fuel regression rate rises with the increase of oxidizer flow rate in zone II at initial moment, and decreases as the oxidizer in zone II increases at middle moment. The average combustion efficiency increases with the increase of oxidizer mass flow rate in zone II when the oxidizer in zone I remains unchanged. Both the average combustion efficiency and the chamber head thermal protection need to be considered in the design of axial injector patterns. Research results of this paper provide an important reference for the design of high performance injector patterns of hybrid rocket motor.

Numerical and experimental investigations on injection effects of orifice injector plate in hybrid rocket motors

Orifice injector plates are widely used in hybrid rocket motors (HRM), but little work has been done on the influences of their design parameters on combustion performances of HRM. The effects of orifice diameter, injection pressure drop and orifices distribution of the orifice injector plate on combustion efficiency of HRM are investigated by experimental and numerical methods in this paper. The hydrogen peroxide (H2O2) and hydroxyl terminated polybutadiene (HTPB) with tube grain are selected as propellant. Firstly, 5 kinds of orifice injector plates with different orifice diameter and orifices distribution are designed, and 7 fire experiments are carried out. The experimental results show that the combustion efficiency of HRM increases with the decrease of orifice diameter and injection pressure drop. Moreover, the combustion efficiency of motor with plat using single-ring distribution is the highest, while the one using three-ring distribution is the lowest. Secondly, 36 orifice injector plates with different design parameters are calculated by three-dimensional numerical simulation of two-phase flow. The liquid oxidizer evaporation, solid-gas interaction, gaseous phase combustion, and turbulence are considered in the simulation model. The conclusions draw from simulation results agree with the experimental data, which illustrates the validity of numerical simulation model. The conclusions based on experimental and numerical results can provide guidance for designing high-performance orifice injector plate in HRMs.

EFFECT OF FUEL PROPERTIES ON THE COMBUSTION OF STORABLE BIPROPELLANTS: ALKANES, ETHANOL WITH HYDROGEN PEROXIDE

Recently, the interest in storable propellants for rocket engines has been driven by the need of cost reducing of space launch: safety, easier handling, and gain in mass structure. Therefore, reduced hazards liquid propellants are considered such as hydrogen peroxide in association with green fuels (kerosene or ethyl alcohol) for propulsion. To this purpose, test benches at lab-scale have been developed at PPrime Institute (Poitiers, France): the one used in the present study is dedicated to the combustion of non-hypergolic propellants at moderate mass-flow rate (ACSEL test bench). The spray is generated by impingement of liquid jets in two configurations: either like doublet or unlike triplet. Both configurations are being studied and implemented in reactive tests with HTP875 (High Test Peroxide of 87.5% by weight) as oxidizer and, respectively, ethanol, iso-octane, and n-decane (as a pure alkane substitute of a kerosene) fuels. The injector plate is implemented in a 105 mm long combustion chamber with different nozzle throat diameters. Tests cover a range of equivalence ratio based on stoichiometry from 0.6 to 2.1 with imposed total mass flow of 15 to 20 g/s. A comparison of the combustion performances and stability is proposed in terms of chamber pressure, total mass-flow rate, equivalence ratio, and characteristic velocity. A focus is given to the effect of the fuel properties on the combustion at steady regime, considering first the effect of the miscibility of the fuel and the oxidizer, and secondly its volatility.

Pressure characteristics of a ram-RDE diffuser

This work focuses on dynamic simulations of a representative ram RDE subsonic diffuser to study ram RDE compatibility with a supersonic inlet. Of particular interest is the propagation of pressure perturbations due to the detonation waves in the RDE combustion chamber. Three-dimensional CFD simulations were completed using a notional diffuser for a ram RDE device operating at 42,000 ft and Mach 2.5. A previous RDE code was used to provide an exhaust forcing function assuming a slot injector plate. A total of 8 different exhaust forcing functions were used, representing RDE air inlet area ratios between 0.3 and 0.6 in single and dual wave operation. Simulations were then completed for the diffuser with the Propel code for each of these exhaust forcing functions. Analysis indicates that the amount of pressure variation seen near the diffuser inlet is very sensitive to the shape of the forcing function at the diffuser exit. The broad pressure rise of the lower area ratios tend to remain intact through the diffuser to the intake with minimal dissipation, while the sharper pressure peak of the higher area ratios tend to be substantially dissipated. For single-wave operation, the broad pressure rise yields pressure variations at the intake that vary from 20.5% (area ratio of 0.3) to 39.1% (a=0.6). For the dual-wave operation, the sharper peak of the pressure rise exhibited even more dissipation. Thus, for dual-wave operation, the amount of pressure variation at the intake varies from 21.0% (a=0.3) to 21.9% (a=0.6) for the different area ratios.

Apollo 11 Command Module Goes on Tour

The exhibit includes Buzz Aldrin's gloves and an injector plate from the rocket's first-stage engine, which was recovered from the bottom of the Atlantic Ocean.

Design Procedure of a Movable Pintle Injector for Liquid Rocket Engines

Variable-area injectors are suitable choices for developing throttleable rocket engines because it is difficult to efficiently control thrust when fixed-area injectors are used. A pintle injector is a variable-area injector that can control the mass flow rate of propellants, thereby replacing a large injector plate having several injector elements with a single injector unit. To achieve proper performance under all throttling conditions, a design procedure was established for a pintle injector based on spray characteristics. The spray angle and Sauter mean diameter of the droplets were critical performance parameters. In spray experiments using water and air as simulants under atmospheric conditions, backlight imaging and laser-diffraction methods were used to measure the spray angle and Sauter mean diameter. The velocity ratio, nondimensionalized opening distance, and Weber number were considered as major nondimensional numbers. Empirical correlations were formulated between performance parameters and nondimensional numbers; these correlations were applied to the design procedure herein. Consequently, a 500 N combustor with a pintle injector was designed using this procedure. A control plan for the pintle-opening distance under lower throttling conditions was also obtained for designing a small combustor. Furthermore, the most sensitive parameters in the injector geometry were confirmed via sensitivity analysis.

Numerical and experimental analysis of heat transfer in injector plate of hydrogen peroxide hybrid rocket motor

This paper is aimed to analyze heat transfer in injector plate of hydrogen peroxide hybrid rocket motor by two-dimensional axisymmetric numerical simulations and full-scale firing tests. Long-time working, which is an advantage of hybrid rocket motor over conventional solid rocket motor, puts forward new challenges for thermal protection. Thermal environments of full-scale hybrid rocket motors designed for long-time firing tests are studied through steady-state coupled numerical simulations of flow field and heat transfer in chamber head. The motor adopts 98% hydrogen peroxide (98HP) oxidizer and hydroxyl-terminated poly-butadiene (HTPB) based fuel as the propellants. Simulation results reveal that flowing liquid 98HP in head oxidizer chamber could cool the injector plate of the motor. The cooling of 98HP is similar to the regenerative cooling in liquid rocket engines. However, the temperature of the 98HP in periphery portion of the head oxidizer chamber is higher than its boiling point. In order to prevent the liquid 98HP from unexpected decomposition, a thermal protection method for chamber head utilizing silica-phenolics annular insulating board is proposed. The simulation results show that the annular insulating board could effectively decrease the temperature of the 98HP in head oxidizer chamber. Besides, the thermal protection method for long-time working hydrogen peroxide hybrid rocket motor is verified through full-scale firing tests. The ablation of the insulating board in oxygen-rich environment is also analyzed.

Measurements and modelling of wall heat fluxes in rocket combustion chamber with porous injector head

A rocket combustion chamber with a porous injector head, which is a new concept for injecting propellants, is tested. The hot-fire tests are carried out using cryogenic oxygen and hydrogen at the mass ratio of oxidizer to fuel of 6 and at pressure of 80 bar. Pressures, wall temperatures, and wall heat fluxes are measured along the axis of the chamber. The wall heat fluxes are determined by the calorimetric method. The wall heat flux has a maximum at a distance of 50–100 mm from the injector plate. To interpret the experimental results, numerical simulations are performed using the commercial CFD code ANSYS CFX. The turbulent flow is modelled by the Favre-averaged Navier–Stokes equations and the shear-stress transport model. The turbulent combustion is modelled using two different models: the extended eddy-dissipation model and the extended coherent flame model. Results obtained with the different models are compared both with experimental data and with each other. The numerical results agree with the experimental data. The extended eddy-dissipation model gives the very good agreement which is achieved by applying additional parameters (‘Flame Temperature’ and ‘Extinction Temperature’) dependent on the mixture fraction.

PRELIMINARY DESIGN OF A HYDROGEN PEROXIDE THRUSTER FOR REMOTE SENSING SATELLITE

In recent years, mono-propellant propulsion has attracted widespread attention in space due to excessive benefits of safety, low cost, simplicity, clean exhaust and high reliability. A 5N monopropellant Hydrogen Peroxide (H2O2) thruster is selected to be utilized with remote sensing (RS) satellites for attitude control and orbit correction. In the present paper, comprehensive theoretical analysis and design parameters are investigated. The effects of change in various design parameters are studied, such as operating pressures (7-16 bar), H2O2 concentration (70-95%) and initial temperature (25-70℃). The crucial design performance of thruster has been carried out by using thermo-chemical software package. A conceptual design was conducted to determine the dimensions of the thruster prototype including the injector plate, catalytic bed, retainer and nozzle part. 70℃ H2O2 preheating temperature might be needed for catalytic bed fast response. The results are described in the form of graphs and tables for the quick and easy analysis.

Combustion Stability Analysis of Rocket Engines with Resonators Based on Nyquist Plots

The paper is concerned with the formulation and validation of an efficient, flexible, and reliable method to assess the influence of resonator rings on combustion dynamics in rocket thrust chambers. The chamber is modeled as a network of low-order acoustic elements that describe three-dimensional waves in cylindrical duct segments with nontrivial boundary conditions and mean flow. The main elements responsible for (thermo)acoustic driving and damping, such as time-lagged fluctuating heat release, injector plate, and short exit nozzle are considered. Particular attention is given to the formulation of a semianalytical resonator ring model, which is capable of handling transverse waves, mode coupling, and mode scattering. A Nyquist plot method based on control theory, which is considerably more efficient than classical root finding techniques, is used to identify the eigenmodes and growth rates of the system. After validating the Nyquist method against analytical results, the stabilizing influence of resonator rings is demonstrated through a representative test case. The key mechanisms induced by the resonator ring that lead to stabilization are also discussed. The proposed method is capable to predict linear stability and give insight in the three-dimensional pressure distribution with very moderate computational effort that allows parametric studies.

Injector Head Transpiration Cooling Coupled with Combustion in Subscale Thrust Chamber

Transpiration cooling coupled with combustion was investigated in an liquid rocket thrust chamber with a transpiration-cooled injector plate. A numerical model was developed using the real gas equation of state. The and combustion process was modeled by the eddy dissipation concept model, which includes the detailed chemical reaction mechanisms in turbulent flows. The permeability and the inertia coefficient of the metal mesh, porous media used for the injector plate were obtained experimentally. The simulation results for the porous plate temperatures and fluxes compare well with hot firing experiments, giving reliable predictions of the combustion, flow, and heat transfer processes in the liquid rocket thrust chamber. The model was also used to investigate the effects of the inlet conditions and the plate material on the transpiration cooling. The results show that locations near the chamber wall and the ignition hole on the plate surface have higher temperatures than other locations. The relatively small thermal conductivity of the porous media is also shown to decrease the plate surface temperature, but increase the thermal stress in the plate.

A feasibility study on using inkjet technology, micropumps, and MEMs as fuel injectors for bipropellant rocket engines

Control over drop size distributions, injection rates, and geometrical distribution of fuel and oxidizer sprays in bi-propellant rocket engines has the potential to produce more efficient, more stable, less polluting rocket engines. This control also offers the potential of an engine that can be throttled, working efficiently over a wide range of output thrusts. Inkjet printing technologies, MEMS fuel atomizers, and piezoelectric injectors similar in concept to those used in diesel engines are considered for their potential to yield a new, more active injection scheme for a rocket engine. Inkjets are found to be unable to pump at sufficient pressures, and have possibly dangerous failure modes. Active injection is found to be feasible if high pressure drop along the injector plate is used. A conceptual design is presented and its basic behavior assessed.

Coupled CFD analysis of hot gas and coolant flow in effusion‐cooled combustion chambers

AbstractThree different coupled CFD analyses of the hot gas and the coolant flow in effusion‐cooled combustion chambers are shown. The common feature of these CFD analyses are the following principles:• solution of the Navier Stokes equations• multi species analysis including hot gas (steam) and coolant (hydrogen)• compressible flow• ideal gas• distributed resistance according to Forchheimer in the combustion chamber wallThe differences between these models are given in the following table:imageThe detailed analysis of the combustion itself is avoided by defining the injector plate of the combustion chamber (left boundaries of the colour plots in figure 2) as an inlet of the by‐product of the combustion (steam). (© 2008 WILEY‐VCH Verlag GmbH & Co. KGaA, Weinheim)

Reactor design optimization based on 3D modeling of nitrides deposition in MOCVD vertical rotating disc reactors

A new modification of the VEECO TurboDisc ® Pioneer reactor (1×2″ wafer) has been designed based on 3D modeling of nitride deposition. The main improvement introduced was for the injector plate design by optimizing the geometrical position of all alkyl zone inlets to provide for optimal control of the growth rate uniformity across the substrate. An optimization technique based on design of experiments (DOE) approach is described which allowed a significant reduction in the amount of calculations required for the optimization. Comparison between trends in effects of rotation rate, chamber pressure and reactant flow as obtained from experimental growth results and modeling computations is reported. It is demonstrated that the use of modeling has reduced the process development time to a few runs and resulted in significant improvement of growth uniformity and alkyl efficiency in the reactor.

96/05246 Fate of metal ions in a brown coal during pyrolysis at low temperatures

This paper is aimed to analyze heat transfer in injector plate of hydrogen peroxide hybrid rocket motor by two-dimensional axisymmetric numerical simulations and full-scale firing tests. Long-time working, which is an advantage of hybrid rocket motor over conventional solid rocket motor, puts forward new challenges for thermal protection. Thermal environments of full-scale hybrid rocket motors designed for long-time firing tests are studied through steady-state coupled numerical simulations of flow field and heat transfer in chamber head. The motor adopts 98% hydrogen peroxide (98HP) oxidizer and hydroxyl-terminated poly-butadiene (HTPB) based fuel as the propellants. Simulation results reveal that flowing liquid 98HP in head oxidizer chamber could cool the injector plate of the motor. The cooling of 98HP is similar to the regenerative cooling in liquid rocket engines. However, the temperature of the 98HP in periphery portion of the head oxidizer chamber is higher than its boiling point. In order to prevent the liquid 98HP from unexpected decomposition, a thermal protection method for chamber head utilizing silica-phenolics annular insulating board is proposed. The simulation results show that the annular insulating board could effectively decrease the temperature of the 98HP in head oxidizer chamber. Besides, the thermal protection method for long-time working hydrogen peroxide hybrid rocket motor is verified through full-scale firing tests. The ablation of the insulating board in oxygen-rich environment is also analyzed.

On reducing the size of liquid separators for injector circulation plate freezers

Liquid separators for injector plate freezers are large because the liquid level rises towards the end of the freezing process. To calculate the volume of liquid being collected in the separator during freezing the freezing time and the heat removed must be evaluated. A simple method of freezing time estimation based on the progression of a phase change front is proposed. The size of separators can be reduced considerably by letting part of the liquid feed by-pass the injector during initial freezing. With this arrangement the injector dimensions are based upon a refrigerating capacity lower than the maximum.