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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]

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]

This paper investigates the role of chaotic analysis and deep learning models in combustion instability predictions. To detect the precursors of impending thermoacoustic instability (TAI) in a swirled combustor with various fuel injection strategies, a data-driven framework is proposed in this study. Based on chaotic analysis, a recurrence matrix derived from combustion system is used in deep learning models, which are able to detect precursors of TAI. More specifically, the ResNet-18 network model is trained to predict the proximity of unstable operation conditions when the combustion system is still stable. The proposed framework achieved state-of-the-art 91.06% accuracy in prediction performance. The framework has potential for practical applications to avoid an unstable operation domain in active combustion control systems and, thus, can offer on-line information on the margin of the combustion instability. [ABSTRACT FROM AUTHOR]

The evaporation characteristics of unsymmetrical dimethylhydrazine droplets with different initial diameters in a subcritical environment were experimentally investigated with the temperature–pressure separation technique. The evaporation processes of unsymmetrical dimethylhydrazine droplets with different initial diameters in this environment have the same general pattern. All the studied droplets exhibit a short transient heating phase and a steady-state evaporation phase obeying d 2 . Notably, the expansion of the transient heating phase gradually increases with increasing ambient pressure. The change in diameter squared ∆ d 2 m a x increases from 1.03% at 1 MPa to 12.48% at 5 MPa. Under subcritical conditions, the evaporation rate decreases linearly with decreasing droplet diameter, and the droplet evaporation lifetime increases linearly. Changes in the initial droplet diameter may still have a large effect on droplets smaller than those studied here. When the ambient pressure is not greater than 3 MPa, the change in the steady-state evaporation time for both medium- and large-diameter droplets accounts for more than 70% of the variation in the droplet evaporation lifetime. As the ambient pressure increases to 4 MPa and 5 MPa, the percentage of the change in the transient heating time contributing to the variation in the droplet evaporation lifetime gradually increases to more than 45%. [ABSTRACT FROM AUTHOR]

This study numerically investigated the thermoacoustic combustion instability characteristics of a scaled rocket combustor based on a hybrid of the Reynolds-averaged Navier–Stokes and large–eddy simulation method. The turbulence–combustion interactions were treated using flamelet generated manifold approach. An unstable case was simulated with detailed reaction mechanisms (GRI-Mech 3.0). The obtained results agree well with experiment data from Purdue University, in terms of pressure oscillations frequency and power spectral density spectrum. The combustion instability mode was identified to be coupled with the first longitudinal acoustic mode of the combustion chamber by dynamic model decomposition method. According to Rayleigh index analysis, the unstable driving source was found to be located near the combustor step, which was further confirmed by time-averaged flow fields. Detailed three-dimensional vortex ring shedding evolutions at the combustor step were tracked with fine time resolution. Results indicate that the combustion instability arises from periodic vortex ring shedding at the combustor step and interacting with the chamber wall. The unburnt reactants were rolled up by the shedding vortex ring, which would not break up until impact with the chamber wall. Therefore, the mixing performance was significantly enhanced, leading to sudden heat release. Consequently, the thermal energy is added to the acoustic field, and the first longitudinal mode is thus reinforced, giving rise to large amplitude axial velocity oscillations which prompt the generation of the new vortex ring. The results of the present investigation will support the design and development of high-performance rocket engines.

A dielectric barrier discharge plasma controlled diffusion flame experimental system was built based on the designed coaxial swirling plasma injector. The air plasma was generated within the annulus gap of the injector by alternating current dielectric barrier discharge. The discharge characteristics and power of plasma injector under different actuation intensities and air flowrates were measured. Through the measurement techniques, such as schlieren imaging, broadband chemiluminescence image and CH* chemiluminescence, the effect and mechanism of plasma on ethylene-air normal diffusive jet and flame was explored. The results showed that a large number of filamentary discharge channels are formed in air plasma. The increase of air flowrates weakened the intensity of discharge to a certain extent. The induced jet generated by the plasma can short the laminar length of the ethylene-air jet, accelerate the transition of the flow jet, enhance the turbulence and the mixing of the fuel and the oxidizer. The higher the actuation intensity, the shorter distance of the cold jet transition zone, the higher the jet turbulence. Depending on the aerodynamic and kinetic effects, plasma can improve the stability of ethylene-air diffusive flame and reduce the lift height between the flame root and injector. The plasma can also expand the flammability limit of ethylene-air flame and make the flame ignited under some conditions that could not be. In addition, the CH* chemiluminescence shows that, in a certain range of discharge voltage, the heat release distribution can be changed on both sides of the flame, and its representative length are generally reduced as the voltage rises. On the contrary, the overtop voltage could lead to a decrease of flame heat release.

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]

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]

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]

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]

The numerical calculation of three-dimensional unsteady combustion for the combustion chamber of LOX/kerosene high pressure staged combustion rocket engine was carried out. By changing the offset ratio of oxygen mass flow rate in the edge area of the injector face, computational studies were conducted to investigate the effects of non-uniform distribution of oxidizer flow on combustion instability for a liquid-propellant rocket engine. The calculation results show that the offset ratio of oxygen mass flow rate changes the distribution of heat release in the combustion chamber. Within a certain range of offset ratio, the non-uniform distribution degree of oxidizer flow enhances the coupling between the pressure and heat release. As a result, it leads to an increase in the pressure oscillation amplitude in the combustion chamber. However, if the offset ratio is too large, the oxygen-fuel ratio will be too small in some regions, which will reduce coupling between the pressure and heat release and increase the damping of combustion instability.

A one-dimensional full transient droplet evaporation model was established under consideration of factors such as high pressure vapour-liquid equilibrium, high-pressure physical property corrections, gas phase dissolution, and shift of interface. The finite volume method was used for discretization to study the migration and evaporation characteristics of the surrogate fuel for kerosene, which was constituting of (mass fraction)80% n-decane and 20%1,2,4-trimethylbenzene, under supercritical conditions. The results show that, under supercritical conditions, the higher the temperature and the pressure, the easier and the sooner the supercritical migration occurs. Before the supercritical migration occurring, there was obvious boundary between the gas and liquid phases. The mass fraction of component was discontinuous, and the gradient of temperature near the interface was large. After the supercritical migration occurring, the surface of the droplet disappeared, there was no obvious boundary between the gas and liquid phases, and the distribution of the components mass fraction and temperature were continuously. With the increase of the initial temperature of the droplet, the time of the supercritical migration was greatly advanced, the rise rate of the droplet surface temperature increased, and the phenomenon of endothermic expansion no longer appeared.

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 air-blower with active magnetic bearings could improve working efficiency and reduce energy consumption by avoiding contact between the rotor shaft and the stator part. The structure and prototype of a magnetically suspended air-blower are herein introduced, and the force models of active magnetic bearings developed. Furthermore, the dynamic models of a rotor shaft with unbalance terms were established to investigate the vibration characteristics of the magnetically suspended air-blower. The vibration characteristics of the rotor shaft with unbalance terms were analyzed, and the complex-field cross-feedback control was designed to suppress the vibration amplitude. Finally, experiments were conducted to verify the theoretical models, and the results indicated that the vibration amplitude of the rotor shaft with unbalance terms could be intensified by the rotating frequency, and the nutation vibration was reduced by 50% through increasing the high-frequency nutation coefficient of the complex-field cross-feedback control model. The results indicated that the vibration analysis of the rotor shaft was meaningful to the design and control of the magnetically suspended air-blower. [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]

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]

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 quasi-DC discharge plasma across the backward facing step of a flame-holding cavity can be used to improve the flame stabilization of the cavity in a scramjet combustor. The effects of plasma filaments on the main flow structures, cavity drag, mass exchange rate, cavity oscillating characteristics, stagnation pressure loss, and combustion efficiency are numerically investigated based on the dominant thermal blocking mechanism. The results show that the cavity shear layer fluctuates dramatically due to the ‘cutting’ effect of plasma, which leads to obvious changes in the fuel mixing and burning in the combustor, and the pressure distribution near the rear edge of the cavity. Some symmetrical structures in the shape of ‘branch pipes’ are observed, which represent the distribution of product water. The periodical producing of them can be attributed to the movement of the cavity shear layer. Compared with the no plasma case, the cavity drag coefficient overall increases 11.3%, yet the mass exchange rate rises sharply over 10 times. The sound pressure level of the cavity representative points is raised by the plasma, and the first dominant frequency of pressure oscillation at the monitor point on the rear edge is twice the plasma actuation frequency. Although the stagnation pressure loss increases, the plasma shows benefits in energy producing that outweigh the energy consumed by the plasma actuator. [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]

Organic gel propellants are promising candidates for a variety of rocket motor and scramjet applications, since they are intrinsically safe and provide high performance. It is well known that organic gel fuel droplets exhibit distinct combustion characteristics compared with conventional liquid fuel droplets, and furthermore an understanding of the ignition delay and lifetime of these droplets is critical to the improvement of combustor design. In this work, investigations of the combustion of unsymmetrical dimethylhydrazine (UDMH) organic gel droplets in different nitrogen tetroxide (NTO) oxidizing atmospheres were conducted using two sets of experimental apparatus. The combustion characteristics under different conditions of temperature and pressure were compared and analyzed based on the flame shapes observed during experimentation. From these trials, an unsteady combustion model was developed and used for the numerical simulation of spray-sized UDMH organic gel droplet combustion in an NTO atmosphere. The hypergolic ignition and burning characteristics of the organic gel droplets under conditions simulating either engine startup or steady state combustion were compared, and changes in ignition delay and droplet lifetime with ambient temperature and pressure were analyzed. The experimental and numerical results show that the UDMH organic gel droplets exhibit periodic swell-burst behavior following the formation of an elastic film at the droplet surface. Each droplet burst results in fuel vapor ejection and flame distortion, the intensity of which declines with increasing ambient pressure. However, the swell-burst period is extended with increasing ambient pressure, which results in potential flameout. Under conditions of low temperature and pressure similar to those at engine startup, the ignition delay and lifetime of spray-sized gel droplets decrease with increasing temperature or pressure, although there is a sharp increase in droplet lifetime when the ambient pressure reaches a critical value associated with flameout. The ignition delay was found to be a rate-limited phenomenon linked to the droplet heating rate. The proportion of ignition delay and droplet lifetime due to droplet heating-up decreased with increasing temperature or decreasing pressure. Conversely, at high temperatures and pressures simulating the engine's steady state operating conditions, the droplets were observed to flameout after several swell-burst periods and both ignition delay and lifetime decreased monotonically with increasing temperature or pressure. The ignition delay time was determined to be rate-limited by gas phase chemical reactions and contributed very little to the overall droplet lifetime compared with the engine startup condition. [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]

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]

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]