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This study presented the numerical simulation of the tangential combustion instability in a hypergolic liquid bipropellant rocket thrust chamber, which applied fuel liquid film cooling method and unlike impinging injectors. The liquid spray was modeled using Lagrangian approach, while the gas was regarded as Euler phase. Stress-blended eddy simulation and finite rate/eddy–dissipation model were adopted to simulate the turbulent combustion process. Consistent with the experiment results, this work successfully simulated the transformation of tangential combustion instability from standing mode to spinning mode. The mean pressure, amplitude and frequency of limit cycle oscillation were in good agreement with the experiment. There was a detailed analysis about the flow field, Rayleigh index, and driving mechanism of the combustion instability. It was found that the oscillation began with hot spots of heat release rate due to the interaction between the spray of impinging injectors and cooling fuel jet. More than that, cooling fuel jet also contributed to drive the oscillation. In the standing mode, injectors in the inner and outer rings drive the oscillation together, while the spinning mode is mainly driven by injectors in the outer ring. The pressure wave is subsonic and its Mach number is close to 1. It was shown that the pressure wave contained a complex structure divided into three parts. This led to the in-phase of the pressure along the axial direction and the double-peak characteristic of the downstream pressure signal. Besides, a positive feedback closed-loop system associated with periodic oxidizer/fuel ratio was believed to sustain the combustion instability. The oscillation can be maintained when pressure, heat release and oxidizer/fuel ratio are coupled together. The analysis results indicate that rotating detonation is an implication to tangential combustion instability. [ABSTRACT FROM AUTHOR]

To suppress the high frequency combustion instability of rocket engines, quasi-direct current (quasi-DC) discharge plasma is considered in combination with special arrangements in a rocket engine combustor. After establishing an internal reacting flow field model of a rocket combustor and a phenomenological plasma model, the influence of quasi-DC discharge plasma on high frequency combustion instability is numerically investigated. The results show that the characteristics of the high frequency combustion instability in the combustor are captured very well through the simulation. Based on the DES turbulence model, a positive feedback process is clearly observed for the self-excited combustion instability. The mean temperature and pressure of the combustor are not sensitive to the occurrence of quasi-DC discharge plasma. The pressure histories of the combustor and their characteristic oscillating frequencies are not affected by the plasma when only one actuator is in operation. However, the oscillating amplitudes of the pressure are reduced within a certain time when several actuators are in operation. The heat release becomes more diffusive and more eddies form around the injector faceplate with an increasing number of plasma filaments. Mainly through the reduction of local heat release, the plasma is capable of affecting the high frequency combustion instability in the combustor. To better control high frequency combustion instability, the optimal arrangement is a double-plasma filaments with symmetrical actuators. • High frequency combustion instability in rocket engine is well captured. • Local heat release is altered by plasma, leading to a weakening in oscillation. • Plasma does not change the phase relationship between pressure and heat release. • Double-plasma actuators scheme is better for combustion instability control. [ABSTRACT FROM AUTHOR]

In order to study the effect of different equivalence ratios on plasma assisted detonation initiation under a rich burn condition, a loosely coupled method was adopted to simulate the detonation formation by alternating current dielectric barrier discharge in a hydrogen–oxygen gas mixture with different equivalence ratios. The spatial and temporal evolution of discharge products were analyzed first. Then, the species distribution, Mach number, thrust wall pressure, and whole history of detonation formation were examined in detail. Results showed that the shapes of the temporal and spatial distribution of discharge products do not change when equivalence ratio varies in one discharge cycle. However, due to the variation of the percentage of fuel and oxidizer, and its impact on the discharge elementary reactions, the number density of every key active particle declines while the decline amplitude decreases when equivalence ratio rises. Although the magnitudes of species concentration are not altered by the plasma, the reacted region extends towards the downstream flow more quickly. The influence of equivalence ratio on plasma assisted detonation initiation becomes remarkable until late in the subsonic stage of the flowfield evolution process in the combustor under a rich burn condition. Furthermore, a larger equivalence ratio leads to a better accelerating effect of the plasma. Through analyzing the dynamic process of thrust wall pressure, it is also found that the DDT process is expedited more notably under a larger equivalence ratio, yet the magnitude of pressure is independent of the plasma. The provision of active particles for the combustion reaction by the plasma and its dilution effect on the fuel are the two intrinsic reasons for the reduction of DDT time and distance and the increase of this reduction degree, when equivalence ratio rises. [ABSTRACT FROM AUTHOR]

A numerical simulation is conducted to characterize the atomization process of like-doublet injector with two different impingement angles. Using thecoupled Level-Set/Volume-of-Fluid (CLSVOF) method for capturing the liquid/gas interfaces, liquid sheet characteristics were efficiently captured, and the simulations were found to agree well with experiment results for both breakup characteristics and Sauter Mean Diameter (D 32 ). This method combines the mass conservation properties of the volume-of-fluid method with the accurate surface reconstruction properties of the Level-Set method for a 3D atomization process simulation of the like-doublet injector. Since the velocity profiles around the stagnation point and impingement point greatly affects the sheet characteristics, the velocity profiles of different axial locations on the cross section for two different impingement angles cases are studied. It shows that the velocity difference caused by the impact of liquid jets is one of the important factors which directly affect the formation of surface wave and the breakup of sheet [ABSTRACT FROM AUTHOR]

Rocket engines are prone to suffer from transverse combustion instabilities, manifesting as periodic oscillations in pressure and heat release rate. Self-excited transverse instabilities in a rectangle multi-injector rocket combustor powered by n-C 12 H 26 and O 2 , are investigated experimentally and numerically in this paper. Transverse combustion instabilities are captured during repeatable experiments, while dynamic characteristics in the combustion chamber and driving mechanisms of instabilities are analyzed via numerical simulation based on stress-blended eddy simulation. Dynamic mode decomposition analysis reveals coupling behaviors between the transverse mode in the combustion chamber and longitudinal mode in the oxidizer post, resulting in propellant mass flow rate oscillations. Substantial evidence demonstrates that the intensities of heat release rate near the end walls are higher than that detected around the chamber center due to its intense coupling to the local first transverse (1W) acoustic mode. Flow field animations indicate strong deflection of oxygen posts and liquid particles by a transverse moving gas mixture leading to flame interactions and collision between adjacent share injectors. The propellant injected into the chamber does not burn immediately and is transported to the combustor end wall. Propellant mixing performance is significantly enhanced during interactions between acoustic pressure and end wall, causing a sudden heat release near 1W pressure anti-node, thereby strengthening pressure oscillations. Finally, a combustion instability driving mechanism associated with propellant mass flow rate oscillations, and interactions between unsteady acoustic pressure and heat release rate is proposed. The insights obtained from this work can aid comprehension of transverse combustion instabilities encountered by rocket engines. • Both liquid fuel and gaseous oxygen are highly deflected by transverse moving gas mixture when transverse combustion instability appearing. • Large flame interactions and collision structures appear inside the combustion chamber. • Combustion instability is driven by pulsated heat release rate near the end walls and periodic supply of propellant mass flow rate. [ABSTRACT FROM AUTHOR]

An experimental study was carried out on the use of a mixture of CO 2 and Xe as the propellant for an electron bombardment ion thruster. The working performance parameters of the ion thruster under different flow mixing conditions were investigated using measurements and diagnostics such as power supply electrical parameter measurements, thrust measurements, and plasma diagnostics. In-depth studies of the particle species and the main electron collision reactions in the discharge chamber were carried out. The results showed that the CO 2 in the discharge chamber led to a significant increase in the mean anode voltage and voltage oscillation amplitude, and the direct participation of CO 2 in the discharge reaction could effectively improve the beam current and thrust of the ion thruster. When 5 sccm of CO 2 was injected into the discharge chamber, the maximum measured thrust was increased by approximately 31.8 % compared to the state without CO 2 injection. The addition of Xe led to a significant decrease in the discharge voltage and hence an increase in the electron number density, which significantly promoted CO 2 dissociation and ionization. The insights obtained from this work can provide ideas for subsequent research related to an ion thruster using a mixed propellant. • A mixed working medium supply method for electron bombardment ion thrusters has been designed. • The ion thruster could start successfully and continuously operate stable under the mixed working medium supply method. • The significant gain of the CO 2 /Xe mixed working medium for the thrust of the ion thruster and its mechanism were verified. • The use of CO 2 as a propellant can effectively improve in-situ resource utilization. • This work may provide more options for the propulsion systems of deep-space exploration missions. [ABSTRACT FROM AUTHOR]

To obtain the mixing enhancement mechanism of H 2 –Water combined jets in supersonic crossflows in a combustor with expanded section for rotating detonation ramjet, the flow field shape and spray structure were studied by experimental and numerical methods. The Eulerian–Lagrangian method was used to investigate the diffusion mechanism and H 2 –Water interaction law of combined jets with different sequences. At the same time, high-speed photography and the schlieren technique were used to capture the flow field. The effects of jet pressure drop, orifice diameter, orifice spacing, incoming Mach number, and other parameters on the penetration depth of water jets were studied. The results of experiment and simulation show that using H 2 –Water combined jets, the penetration depth of the jet spray can be greatly increased and the jet mixing effect can be significantly improved, which will contribute to the engine's ignition and stable combustion. In the case of pre-water/post-H 2 , the penetration depth of the hydrogen jet is greater. In the case of pre-H 2 /post-water, the hydrogen jet raises the water spray mainly by protecting the integrity of the water column. • The mechanism of the mixing process of H2-Water combined jets in supersonic crossflows was investigated. • Eulerian-Lagrangian method was used to investigate the H2-Water interaction law in a combustor with an expanded section. • The penetration depth of water jets under different parameters was studied experimentally. [ABSTRACT FROM AUTHOR]

The detonation wave in a rotating detonation engine is highly adaptable to the incoming flow, making the wave easier to control. In this study, a numerical simulation method is used to analyze the working process and flow field structure of a rotating detonation model engine with dual cavity injection of an H 2 /air mixture by controlling the injection pressure ratio of the dual cavity and the number of detonation wave heads. It is found that the rotating detonation engine offers the possibility to control the thrust vector with two different modes. The first is a one-cycle alternate control mode with a small injection pressure ratio. Here two deflections occur in different directions occur across one detonation wave propagation cycle, but the overall deflection direction is in the low-pressure region. The second is a one-way control mode, with a large injection pressure ratio, and the deflection direction towards the low-pressure region. For the multi wave-mode, it belongs to one-way control mode because of constant deflection direction in the low-pressure area. From the perspective of thrust distribution along the circumference, the one-way control strategy satisfies the ability of a rotating detonation thrust vector control. • RDE theoretically owns the ability of the thrust vector control with two modes. • Analyze the specific impulse and thrust of the rotating detonation engine during the processes of the thrust vector control. • One-cycle alternate and one-way control modes have different processes, but the latter is more suitable for RDE applications. • Numerical simulation with dual cavity injection was adopted to analyze the TVC working process and flow field structure. [ABSTRACT FROM AUTHOR]

This study investigates the atomization process of a liquid jet in supersonic crossflow at a Mach number of 2.1 by using a multi-scale simulation method. This hybrid numerical method can realize accurate simulation of the primary and secondary breakup of the liquid jet with large scale instabilities and small droplet dynamics under compressible supersonic crossflow. The Eulerian framework is applied for capturing the gas-liquid interface by a coupled Level set and Volume-of-fluid (CLSVOF) method. Based on the Delayed Detached Eddy Simulation (DDES), the adaptive mesh refinement (AMR) method is introduced to predict the breakup of liquid column, blocks and filaments. The Lagrangian framework is applied for the dispersed droplets generated under the conversion criteria during the atomization process. The hybrid CLSVOF-LPT method can greatly reduce the loss of liquid phase mass and save computing resources. The distribution morphology of droplets in a cross-sectional plane from the simulation agreed well with the experimental data. Besides, the trends of the jet trajectory and primary breakup behavior of liquid column are well captured. • The Hybrid CLSVOF-LPT numerical method is applied in the atomization of a liquid jet in supersonic crossflow. • The flow field (including shock waves and separation zones) and the spatial distribution of the spray are well captured. • The 'Ω' shape of the cross-sectional spray distribution is in good agreement with the experimental data. • Sprays are found in the separation zone, due to the separated recirculating vortex and its reattachment. [ABSTRACT FROM AUTHOR]

The diffusion process and spatial distribution of a liquid jet in supersonic crossflow (Mach number of 2) in a model combustor with an expanded section were investigated on the basis of experiments and simulations. Experiments were performed using high-speed photography, high-speed shadowgraph and laser-based particle size analyzer to investigate the spray structures, flow fields and droplet diameter, respectively. The Eulerian-Lagrangian method coupled with the Kelvin-Helmholtz/Rayleigh-Taylor breakup model were used in the simulations. Moreover, delayed detached eddy simulation numerical method based on the two-equation shear stress transport turbulence model was applied. The spray distributions and flow fields predicted in the simulation, including the spray penetration, cross-sectional distribution, droplet size in the far-field, the bow shock, large-scale vortices and recirculation zones were in good agreement with experiments. The evolution of spray in supersonic crossflow could be divided into three stages, aerodynamic induced liquid column fracturing, expansion wave promoted spray accelerating, and compression wave disturbed spray blending. Overall, the channel configuration (with expended section) changed the direction and strength of the supersonic airflow. The force of the crossflow determined the spray diffusion. • The diffusion process and spatial distribution of LJISC in a combustor with an expanded section were investigated. • The simulation results such as spray structures, airflow structures were comprehensively verified experimentally. • Typical structures of LJISC in a rectangular channel with an expaned section could be divided into three regions. [ABSTRACT FROM AUTHOR]

In view of the application requirements for accurately measuring the micro thrust of ion thrusters in space electric propulsion, an experimental study of micro thrust measurement under variable working conditions was carried out based on an 8 cm ion thruster. By combining highly sensitive thrust response parameters such as gas flow, anode voltage and grid voltage, the ion thruster can achieve high-precision continuous thrust adjustment in the range of 1–18.6 mN. The calibration results of the thrust measuring device were compared in the cold state and the stable state of thermal equilibrium of the ion thruster, respectively, and the measured thrust was compared with the thrust value calculated based on beam current characteristic measurement. The experimental results showed that compared with the cold calibration, the calibration results of the thrust measuring device in the stable state of thermal equilibrium of the ion thruster are more suitable for measuring the thrust of the ion thruster, and the measured thrust value is in good agreement with the thrust value calculated based on beam current characteristic measurement, and the relative error of the two within the research range is less than 4%. • The 8 cm ion thruster can achieve high-precision continuous thrust adjustment in a wide range of 1–18.6 mN. • Thermal calibration can reduce thrust measurement errors effectively. • The measured thrust is in good agreement with the thrust value calculated based on beam current characteristic measurement. [ABSTRACT FROM AUTHOR]

Abstract Based on nano-tracer based planar laser scattering (NPLS) and background oriented schlieren (BOS) based wavefront sensing (BOS-WS), the fine structures and aero-optical effects of Mach 3.4 flow over the backward facing step with or without vortex generators (VGs) control were investigated experimentally. The structures were revealed in the instantaneous and time-averaged flow images with high spatial–temporal resolution. The shape, size and angular orientation of the large-scale structures is characterized by two-point spatial correlations method. The BOS results displayed the aero-optical effects distorted by the flow field with or without VGs control. It is concluded that the recirculation region in the case with VGs control is smaller based on the comparison of the optical path difference (OPD) distributions using BOS-WS. The OPD will be larger downstream the reattachment shockwave than in the case without control. The root-mean-square value of OPD of the observation field in the case without control is 0.7317 λ , while with VGs control is 0.5388 λ , where λ is the incident light wavelength. [ABSTRACT FROM AUTHOR]

This study investigates the mixing processes of liquid jets in supersonic crossflows at a Mach number of 2.1 by using numerical simulations that employ the Coupled Level-Set/Volume-of-Fluid (CLSVOF) method, which is based on the Delayed Detached Eddy Simulation (DDES). The instantaneous spray morphology from the simulations agreed well with the experimental data. The spray presents the 'Ω' shape when viewed from the front, and the maximum error in the penetration depth was 15% from a side perspective. The direction of the air movement determined the shape of the liquid column shearing and breaking. In turn, the breaking process of the liquid jet significantly affects the structure of the air flow. Three-dimensional flow around the liquid column occurred when the supersonic incoming flow impinged on the liquid column, ultimately forming counter-rotating vortex pairs (CVP) in the inner region of the spray. Evolution of the reverse vortex pair accelerated the mixing process of the spray. Single liquid jet was used as a basis to further study double jets in Mach 2.1 airflow. The results show that the breaking length and penetration depth of the downstream jet are significantly improved. [ABSTRACT FROM AUTHOR]

Highlights: [•] Oxidant convective conditions affect the burning behavior of gel UDMH droplets. [•] The enveloped flame is converted to the escaped flame as convective velocity rises. [•] The enveloped flame is more helpful to the burning process. [•] Microexplosion intensity of enveloped flame rises with convective temperature. [•] Microexplosions intensity of escaped flame is higher, but frequency was lower. [Copyright &y& Elsevier]

Self-excited longitudinal combustion instabilities were analyzed in a sub-scale combustor operating with a propellant combination of O 2 /CH 4. Couplings between acoustics, hydrodynamics, and heat release were investigated numerically based on an experimental setup. The numerical results were in good agreement with the experimental data. Various post-processing methods have been adopted to explain the thermoacoustic behavior and reveal the underlying unstable combustion mechanisms. The classical longitudinal acoustic characteristics of the combustion chamber were identified by the pressure mode shape and dynamic mode decomposition (DMD) analysis. It was found that the incoming streams along the injector were periodically hindered by the traveling pressure wave originating from the chamber, thus, resulting in an oscillatory propellant supply and a pulsating heat release. This, in turn, feeds energy into the acoustic pressure oscillations, and maintains the combustion instability. Meanwhile, the coherent vortex structures in the reacting mixing-shear layer were found to be the principal reason for sustaining combustion instabilities. Besides, a positive feedback closed-loop system associated with periodic vortex shedding and merging is believed to sustain combustion instabilities. The plot of the Rayleigh index further confirmed that the shear layer region drives thermoacoustic instabilities. [ABSTRACT FROM AUTHOR]

Self-pulsation characteristics of a liquid-centered swirl coaxial (LCSC) injector are investigated numerically based on the coupled volume of fluid VOF and Level Set (CLSVOF) model. Good agreements of simulation results and experiment data, in aspect for self-pulsation frequency and spray angle, are generally achieved. On the basis of the flow dynamics of the injector, mechanism of spay self-pulsation in recessed LCSC injector is proposed. The self-pulsation arises from the periodic blocking of annular gas by the liquid sheet. Self-pulsated "Christmas tree" spray patterns result from the periodic vortex rings generating and shedding around the liquid sheet. Theoretical analysis demonstrates that the self-pulsation frequency is almost the same as the frequency of the liquid sheet impacting the recess chamber wall, and the model may be able to predict the self-pulsation frequency. Pressure oscillation inside the air-core is closely related to the primary breakup of the liquid sheet, and the frequency is consistent with the self-pulsation frequency. Effects of the gas-liquid mass flow rate ratio (GLR) on the self-pulsation frequency and pressure oscillation frequency inside the air-core are negligible. The self-pulsation intensity peaks under conditions with the critical mixing flow pattern, while, the pressure oscillation intensity inside the air-core increases with the elevating of GLR. It seems that the spray self-pulsation could improve the spray angle. [ABSTRACT FROM AUTHOR]

The self-pulsation characteristics of a liquid-centered swirl coaxial (LCSC) injector are investigated numerically based on the coupled level set and volume of fluid (CLSVOF) model. Effects of recess length on self-pulsation characteristics are evaluated and discussed. The results demonstrate the crucial role of recess length in exciting self-pulsation. As the recess length increases, spray patterns vary from stable hollow cone to self-pulsated "Christmas tree" shape, and the characteristic frequency increases marginally. Compared with the stable spray, when self-pulsation occurs, the droplets are distributed more widely in the radial direction. Self-pulsation is easily excited under conditions with the critical mixing flow. Several methods to suppress self-pulsation have been proposed. The breakup of the liquid sheet is successively dominated by surface waves and spray self-pulsation as the recess length increases. The waves in both the azimuthal and the axial directions are captured and investigated. The unstable waves are composed of sinuous mode and varicose mode. The varicose waves are responsible for the primary breakup of liquid sheet. The results for the self-pulsated spray suggest that the primary breakup of the liquid sheet is coupled with pressure oscillation inside the confined region. Self-pulsation of recessed LCSC injector is suggested to be caused by the periodical blocking of the annular gap by the liquid sheet. [ABSTRACT FROM AUTHOR]