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The feasibility of using preheated kerosene and pre-evaporation to assist in liquid kerosene detonation ignition and self-sustaining combustion under near-ramjet limit conditions was investigated. Additionally, the detonation decoupling and self-sustaining mechanisms were analyzed in depth. Owing to the forced heat transfer effect of supersonic flow on kerosene droplets, self-sustaining detonation combustion could not be achieved using preheated kerosene alone. The detonation wave was still decoupled due to the forced heat exchange effect of the cold kerosene droplets on the detonation wave, even if the distribution of C12H23 reached the detonation conditions through the pre-evaporation of fuel. [ABSTRACT FROM AUTHOR]

To evaluate the nanosecond pulsed discharge plasma-assisted evaporation and ignition attributes of hydrocarbon fuel droplets, numerical models are established. The influences of nonequilibrium plasma on the characteristics of n-decane droplet evaporation and ignition are numerically studied through a loosely coupled method employing a one-dimensional (1D) fully transient model of droplet evaporation and ignition and a model of nanosecond pulse plasma discharge. The results show that the reaction rate of the initial phase of n-decane droplet ignition is increased by nonequilibrium plasma. In addition, both the evaporation and ignition of the droplet are accelerated. Increasing the reduced electric field of the discharge leads to decreases in the droplet ignition delay time, survival time, and shortening amplitude. The evaporative cooling effect, which occurs at the initial phase of droplet ignition and typically decreases the local temperature surrounding a droplet surface, is weakened by the plasma. As the reduced electric field increases, the time to generate a high-temperature zone (>1800 K) decreases, while its duration increases. The initial phase of n-decane droplet ignition, in which the flame temperature increases while the flame radius decreases, is strongly affected by the plasma during the initial ignition phase. However, the plasma has little impact on the peak flame temperature and the flame radius during the stable combustion phase of the droplet. According to reaction kinetics analysis, the plasma directly interferes with the elementary reactions R330 (nC10H22 + O = 3-C10H21 + OH) and R331 (nC10H22 + O = 2-C10H21 + OH) of n-decane/air combustion. Moreover, the dissociation and oxidation processes of intermediate products are accelerated. Then, the important reaction rates, which determine the ignition delay time, increase indirectly. Thus, the overall ignition delay time of the n-decane droplet is shortened. [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 plasma filament is generated upstream of the fuel jet through quasi-dc discharge in a scramjet combustor. The effects of the plasma filament on fuel mixing and combustion are numerically investigated based on the dominant thermal blocking mechanism. Results have shown that temperature distribution of the whole flowfield is changed prominently due to heat release from the plasma filament. Since the shock waves induced by the transverse fuel jet are partly weakened and move upstream by the plasma filament, the separation zone upstream of the fuel orifice can be controlled to provide a better combustion environment. Compared with the no-plasma case, more water appears in higher space above the bottom wall while less near the bottom wall of combustor. Although the stagnation pressure loss increases a little owing to the plasma filament, the combustion efficiency rises obviously with only a little external energy addition. [ABSTRACT FROM PUBLISHER]

Effects of secondary gas injection on flame shape and combustion stability of a premixed swirl methane/air flame were investigated experimentally. Without secondary gas injection, an attached M-shaped flame was observed with pressure oscillations up to a sound pressure level of 68 dB. The flame exhibited significant oscillations due to flame–vortex interactions induced by velocity fluctuations. The secondary methane or air was separately injected into the outer recirculation zone to suppress the flame oscillations at a volume flow rate of 0.5% of the primary flow rates. For lower primary air flow rates ( Q air ≤ 60 L/min), both secondary methane and air injections were able to suppress the thermoacoustic instability and performed the sound pressure level reduction of 45 dB. Meanwhile, the flame transformed from an attached unstable M-shaped flame to a detached stable V-shaped flame. The secondary methane or air injection served to stabilize the flame and anchor its spatial heat release distribution by limiting the variation of flame shape in the outer recirculation zone, suppressing thermoacoustic oscillation. For higher primary air flow rates ( Q air = 80 L/min), the secondary methane injection could reduce sound pressure levels to noise levels, while secondary air injection could not. The velocity disturbances and heat release fluctuations introduced by secondary methane injection could lead to a transformation in flame shape and achieve the suppression of thermoacoustic oscillation. However, the velocity disturbances induced by secondary air injection were insufficient to stabilize the flame. [ABSTRACT FROM AUTHOR]

The response characteristics of the bi-swirl coaxial injector under oscillating backpressure were investigated numerically based on coupled level-set and volume of fluid model. The backpressure oscillation frequency was set ranging from 1000 to 5000 Hz. The frequency and amplitude response characteristics of liquid films were analyzed under different injection conditions. The amplitude of the pressure response inside the injector was determined by the acoustic damping of the system. Due to the central gas core inside the inner injector, the inner injector manifold was found to be more sensitive to the backpressure oscillation. The frequency characteristic of the liquid film response indicated the presence of the characteristic frequency where the response amplitude was maximum. This characteristic frequency was linearly positively correlated with the mass flow rate of the inner injector, while the flow in the outer injector could suppress the response amplitude of liquid films. Meanwhile, the nonlinearity of the response was dominated by the inner liquid film and was also the strongest at the frequency with the highest response amplitude. The simulated flow fields showed that the oscillation of the ambient airflow periodically hindered the liquid film, leading to Klystron effect. Finally, it was suggested that the characteristic frequency was related to the axial velocity of the inner liquid film at the injector outlet and the acoustic length of the gas core. In addition, the response amplitude of liquid films was found to be controlled by the spray cone angle and the total mass flow rate of propellants. [ABSTRACT FROM AUTHOR]

Thermoacoustic instability (TAI) presents a critical challenge for lean-burning combustors and rocket engines. The early detection of instability is crucial, and to address this, a data-driven prediction framework has been established for TAI in a sub-scale rocket combustor with variable chamber length. Nonlinear combustion features are generated from time series of dynamic pressure using recurrence matrices. Deep learning models are then utilized to train these features and predict the proximity of impending TAI. The performance of the proposed method is investigated through cross-validations of 12 groups of hot-fire test datasets. Remarkably, the prediction performances are in good agreement with measured experimental data, with most instabilities being predicted dozens of milliseconds in advance. This capability paves the way for the early implementation of active control systems in full-scale combustors in the future. The prediction performances are also compared with state-of-the-art TAI prediction methods. [ABSTRACT FROM AUTHOR]

In this study, the temperature–pressure separation technique was used to successfully obtain the process of the unsymmetrical dimethylhydrazine droplet diameter variation with time during evaporation in subcritical, supercritical, and far-supercritical environments at temperatures and pressures of 473–773 K and 1–9 MPa, respectively. The results show that unsymmetrical dimethylhydrazine droplet evaporation has two characteristic stages, transient heating and steady-state evaporation, under low-pressure and low-temperature environments; the steady-state evaporation stage influences droplet evaporation life more than the transient heating stage. Under high-pressure and high-temperature environments, droplet evaporation shows only a single feature of steady-state evaporation, and the droplet expansion process in the transient heating stage disappears. The normalized droplet lifetime decreased exponentially with temperature and decreased linearly and monotonically with pressure. The evaporation rate constant increases linearly and monotonically with temperature and exponentially with pressure. [ABSTRACT FROM AUTHOR]

Compared to the continuous mode of a high-frequency alternating current (HFAC) source, the burst mode is believed to be capable of controlling flow and combustion while saving discharge energy. Thus, the effect of burst mode on plasma-assisted flame stabilization, which is driven by an HFAC source, is investigated based on a plasma manipulation diffusion flame experiment system. A specially designed injector with a coaxial dielectric barrier discharge configuration is used to generate a methane-air diffusion flame and air plasma. The influence of the actuating parameters of the burst mode on plasma flame stabilization and the discharge power consumption are studied. The results show that the plasma injector has the typical volt–ampere characteristics of dielectric barrier discharge. The discharge clearly exhibits regular unsteady operation characteristics when the source is in burst mode. Two flame stability thresholds are found with increasing discharge voltage. According to the first and second flame stability threshold values and the flame pattern, raising the control frequency and the duty cycle is favorable for flame stabilization. However, for a fixed duty cycle of 50%, because the time scale of the actuator-off duration is longer than that of combustion, when the control frequency is too low, the flame cannot be stabilized even though the applied voltage is very high. For a fixed control frequency of 100 Hz, the flame is also difficult to stabilize when the duty cycle is too small. Mainly due to the unsteady gasdynamic effects of plasma in the burst mode, this approach can achieve the same flame stabilization effect as that of the continuous mode while reducing the energy consumption. The cost-effectiveness ratios of the two optimum actuation schemes are only 6.4% and 10.1%. In addition, the burst mode is better than the continuous mode in avoiding undesirable arc discharge. To better stabilize the flame while maintaining a relatively low cost-effectiveness ratio, actuator setting schemes with high frequency and a small duty cycle should be selected. [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]

Automated target species selection dynamic adaptive chemistry (automated DAC) based on the directed relation graph with error propagation is a dynamic mechanism reduction method that can alleviate the huge computational overhead of high-precision large eddy simulations. However, the accuracy of this algorithm is often assessed based on the temperature and ignition delay time, which are a posteriori results and do not reflect the physicochemical nature. Because of this, we proposed a new criterion based on the tangential stretching rate (TSR) value to study the automated DAC method from the perspective of chemical reactions, which can characterize the most energetic reaction mode. First, we tested the new criterion by applying it to the analyses of the simulations adopting the detailed mechanism and automated DAC. The results verified the performance of the new criterion and showed that a new TSR value oscillation error phenomenon was discovered by it. This error is caused by the inappropriate cut-off of the important species and chemical reactions in the DAC-generate skeleton mechanisms. Second, we studied the measures to eliminate this newly emerged error. We found that reducing the search threshold, increasing the number of target species of the automated DAC method, and modifying the automated DAC to the time-correlated automated DAC by adding a time correlation term can alleviate this TSR value oscillation error. In addition, constant pressure autoignition simulations of methane/air were implemented to support this study. The TSR-computational singular perturbation participation indices were also introduced for assistance. [ABSTRACT FROM AUTHOR]

A planar Langmuir probe, which can be used for both ion electric propulsion beam characteristics measurement and plasma diagnosis, was developed and applied to plume measurement and diagnosis experiments of an 8 cm Kaufman ion thruster to address the growing demand for miniaturized ion thruster plume measurement and diagnosis. The experimental results show that the plume area exhibits good symmetry along the central axis, and the beam dispersion angle of the 8 cm ion thruster is approximately 22.7°. The radial distributions of the floating potential, plasma space potential and ion density are similar, showing a high distribution in the middle and a relatively low distribution at the two ends. The miniaturized ion thruster plume measurement and diagnosis system is not only easy to manipulate, but also fast and efficient for plasma plume measurement and diagnosis, which provides more complete data support for analyzing the working state of the miniaturized ion thrusters and optimizing the design of the thruster structure. [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]

A methane/oxygen mixture is considered to be an appropriate propellant for many future rocket engines due to its practicality and low cost. To better understand the combustion instability in methane/oxygen-fed rocket engines, the spontaneous transverse combustion instability in a rectangular multi-element combustor (RMC) was analyzed both experimentally and numerically. Severe combustion instabilities occurred in the RMC during repeatable hot-fire tests. The physical mechanisms were systematically investigated through numerical simulations based on the stress-blended eddy simulation and flamelet-generated manifolds method with detailed chemical mechanisms (GRI Mech 3.0). The numerical results for the frequency spectrum and spatial modes agree well with the theoretical analysis and experimental data. The driven regions of the combustion instability were identified on both sides of the combustion chamber through a Rayleigh index analysis. The longitudinal pressure oscillations in the oxidizer post were found to be coupled with the transverse pressure waves in the combustion chamber and led to periodic oscillations of the mass flow rate of propellant. Moreover, the mixing was highly enhanced when the pressure wave interacted with walls of the combustion chamber. Therefore, a sudden release of heat occurred. The pressure oscillations were enhanced by pulsated heat release. A closed-loop system with positive feedback associated with periodic oscillations mass flow rate of the propellant, and sudden heat release, was believed to account for the present combustion instability. [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]

This paper presents an experimental study on rotating detonations in a hollow combustor with the slit-orifice nozzle. The experimental results reveal that the propagation speed of detonation waves increases with the rise of mass flow rates and is greater than the Chapman–Jouguet detonation speed (VC−J). Furthermore, numerical simulations of rotating detonation in a non-premixed three-dimensional cylindrical combustor have been conducted based on a multispecies reacting code. The influence of two mixing schemes—that is, slit-orifice and coaxial injector—on detonation waves are studied to determine whether the characteristics of detonation waves tend toward high-frequency combustion instability due to changes in the mixing scheme. It is found that the slit-orifice scheme's detonation speed, pressure, and temperature are significantly higher than those of the coaxial injector scheme. In particular, the detonation speed of the former reaches 124% of the VC−J, while that of the latter is only 80.5% of the theoretical value. The numerical results reveal that the low-speed detonation is caused by the deterioration of the hydrogen (H2)/air mixing conditions. Moreover, the flow-field structures of two mixing schemes were comparable, both containing transverse detonation waves, oblique shocks, contact surfaces, and wedge-shaped reactant regions. Furthermore, the Rayleigh index analysis showed that the unsteady heat release was in phase with the pressure fluctuations, amplifying the pressure. Therefore, it is suggested that high-frequency combustion instability may be a manifestation of rotating detonation waves under poor mixing conditions. [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]

Self-excited combustion instabilities of transverse modes were experimentally investigated in a rectangular multi-injector model combustor, operating with the bipropellants O2/CH4. The propellants were injected through a linear array of five oxidizer-centered shear coaxial injectors into the combustor. High-amplitude limit cycles obtained in hot-fire tests were analyzed in detail. Different combustion instability modes, including first and second width modes, were observed in cases with three different injection distribution schemes. Hence, the injection distribution strongly determined the combustion dynamics. One insight can be gained that the stable combustion could be achieved by properly designing the propellants' injection distributions. [ABSTRACT FROM AUTHOR]

The process of dielectric barrier discharge (DBD) plasma flow-separation control on screw propeller is numerically simulated based on the unsteady compressible Navier-Stokes equations loosely with DBD-plasma body force aerodynamic actuation mechanism. The effect and the mechanism of plasma flow control to enhance the propeller aerodynamic characteristics are discussed. The results show that the propeller aerodynamic characteristics will decrease as the altitude increase without plasma flow control, but it will first increase and then decrease as the altitude increases with plasma flow control. The effect of DBD-plasma flow control on the distribution of pressure is small at ground condition, and the difference of pressure distribution between general propeller and plasma efficiency-enhance propeller will be increased as the altitude increases. The effect of DBD-plasma flow-separation control will increase as the altitude increases. Mainly, the mechanism of the plasma flow control to enhance the propeller aerodynamic is obtained as follows. The flow separation on the propeller blade surface could be effectually controlled by DBD-plasma aerodynamic actuation, thereby increasing the efficiency of the propeller. The pressure of the windward surface of the blade will decrease by the plasma body force, which will increase the thrust of the propeller. [ABSTRACT FROM PUBLISHER]

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]

Lift-enhancement–drag-reduction technology is strongly required by near-space vehicles with low Reynolds number. It is known that a flow control method by a surface dielectric barrier discharge (SDBD) plasma can play an important role in this field. In order to obtain the discharge characteristics and evaluate the flow control effect of a SDBD actuator, the nanosecond-pulse discharge and induced flow field by the SDBD plasma are simulated at various altitudes using discharge-aerodynamics models. The results show that the ignition voltage decreases with altitude and it is very easy to discharge in near space. Compared with a SDBD at ground level, the plasma is produced on both sides of the exposed electrode and distributes more uniformly in near space. Although the body force generated by the SDBD actuator is less effective in inducing a jet with nanosecond-pulse excitation than that with alternating voltage excitation, the induced jet by body force is with longer extent, thicker profile and higher velocity in near space than at ground level. The plasma bulk heating should be taken into account for nanosecond-pulse excitation. The Joule heating of electrons is the main source of plasma bulk heating which acts as a micro-explosion and mainly induces pressure perturbation. The discharge at ground level is like a ‘point explosion’, but a ‘region explosion’ in near space, which indicates a diffuse distribution of energy, should be responsible for the fact that the effect of bulk heating is reduced in near space. [ABSTRACT FROM AUTHOR]

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]