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Fischer-Tropsch (F-T) synthesis is an important route to achieve the clean fuel production. The performance of gas–liquid separation equipment involving in the progressive condensation and separation of light and heavy hydrocarbons in the oil-gas products has become a bottleneck restricting the smooth operation of the F-T process. In order to remove the bottleneck, a gas–liquid vortex separator with simple structure, low pressure drop and big separation capacity was designed to achieve the efficient separation between gas and droplets for a long period. The RSM (Reynolds Stress Model) and DPM (Discrete Phase Method) are employed to simulate the flow characteristics and liquid distribution in the separator. The results show that the separation efficiency is influenced by the flow field and liquid phase concentration in the annular zone. The transverse vortex at the top of spiral arm entrains the droplets with small diameter into the upper annular zone. The entrained droplets rotate upward at an angle of about 37.4°. The screw pitch between neighbor liquid threads is about 0.3 m. There is a top liquid ring in the top of annular zone, where the higher is the liquid phase concentration, the lower is the separation efficiency. It is found that by changing the operating condition and the annular zone height the vortex can be strengthened but not enlarged by the inlet velocity. The screw pitch is not affected by both inlet velocity and annular zone height. The liquid phase concentration in the top liquid ring decreases with both the increases of inlet velocity and annular zone height. The total pressure drop is almost not affected by the annular zone height but is obviously affected by the inlet velocity. When the height of annular zone is more than 940 mm, the separation efficiency is not changed. Therefore, the annular zone height of 940 mm is thought to be the most economical design.

An analysis has been completed for a comprehensive set of vertical, dilute-phase pneumatic conveying pressure drop data from an investigation by Flatow. The data were collected in the established flow region for eight different materials conveyed in 0.05-, 0.10-, 0.20-m internal diameter, 20-m tall steel risers. Particle velocities derived from the pressure drop data were used to develop an equation of motion that includes terms for pipe diameter, terminal velocity, coefficient of restitution, and particle shape. The best data fit was achieved using the actual gas density and the actual gas velocity adjusted for voidage. Adjusting the terminal velocity for voidage, an approach recommended by many investigators, did not improve the fit for reasons identified by the present research. Using the equation of motion, particle velocities were predicted and used to calculate total pressure drops that are within ±15% of the measured values. The calculated values also produce the characteristic trough-shaped total pressure drop curves allowing the minimum pressure drop gas velocity to be determined without recourse to a separate correlation. A comparison with other studies using shorter risers indicates that data from these studies likely include acceleration effects. A separate study will investigate this observation further.

Multi-jet interactions are characterized by a number of shock and expansion waves, formation of turbulent shear layers and subsonic recirculation zones. The problem of the occurrence of unsteady, transient and oscillatory, processes arising in the nozzle devices of rocket engines at different flight modes is discussed. The main flow regimes, the sequence of their change and qualitative patterns of shock-wave structures that arise in each of the regimes are considered. The main emphasis is paid to the causes of occurrence and mechanisms of maintenance of low-frequency oscillations of the base pressure. The regularities of changes in the amplitude-frequency characteristics of oscillations are studied depending on the main parameters of the problem. Areas of existence of oscillatory processes are found, and a search is carried out for the ranges of design parameters, geometric characteristics of nozzle assemblies that provide a non oscillating flow of gas-dynamic processes. Hysteresis phenomena of shock-wave structure at increase or decrease in the reservoir total pressure is observed.

A common objective in the analysis of turbomachinery components (nozzle guide vanes (NGVs) or rotor blades, for example) is to calculate performance parameters, such as total pressure or kinetic energy (KE) loss coefficients, from measurements in a nonuniform flow-field. These performance parameters can be represented in a range of ways. For example, line-averages used to compare performance between different radial sections of a 3D component; plane-averages used to assess flow (perhaps loss coefficient) development between different axial planes; and fully mixed-out values used to determine the total loss associated with a component. In the literature, the weighting method used for line- and plane-averaging (e.g., area, volume flow, mass flow, or entropy-flux) is sometimes regarded as an unimportant issue. Indeed, many authors neglect to even state which weighting method was used in their work. In certain low-speed test cases, or where measurements are made a long distance from the component, the nonuniformity in the flow will be relatively small and the practical difference between different weighting methods may be negligible. However, in high-speed applications or for measurements close to a component trailing edge, this becomes increasingly unlikely. In this paper, we compare a range of methods for calculating aerodynamic performance parameters—for example, the kinetic energy loss coefficient—including plane-average methods with different weighting schemes and several mixed-out methods. We analyze the sensitivities of the different methods to the axial location of the measurement plane, the radial averaging range, and the exit Mach number. We use high-fidelity experimental data taken in several axial planes downstream of a cascade of engine parts: high-pressure (HP) turbine NGVs operating at transonic Mach number. The experimental data are complemented by computational fluid dynamics (CFD). We discuss the underlying physical mechanisms which give rise to the observed sensitivities. The objective is to provide guidance on the accuracy of each method in a relevant, practical application.

The understanding of the flow characteristics and effect of gas-solid interactions in pneumatic risers is fundamental to investigate to ensure effective design cost-effective operation. Thus, to understand the effect of gas-solid interactions on the hydrodynamics of newly proposed conversing risers, this study mainly focused on predicting pressure drop in the dilute phase pneumatic conveying system. The experiments were conducted in a converging riser having a convergence angle of 0.2693°. Various solid particles such as sago, black mustard, and alumina have been considered to study the effect of particle sizes and density on the pressure drop. The experimental outcomes indicate that the total pressure drop increases with an increase in the solid density and gas mass flow rate. Moreover, smaller particle sizes are also increased the pressure drop. An empirical correlation is developed for the prediction of total pressure drop ΔPT in converging pneumatic riser via dimensional analysis. All dependent variables such as particle and air density, drag force, acceleration due to gravity, the mass flow rate of air and particle, the diameter of particle and converging riser, the height of converging riser were considered to develop the empirical correlation. The established relationship is tested, and experimental data have been fitted for its validation. The estimated relative error of less than 0.05 proved the significance of the developed correlation. Hence, it can be stated that the established relationship is useful in studying the effects of various parameters on the pressure drop across the length of the conversing riser.

The present study modifies the structural design of a shell-and-tube heat exchanger by considering two key parameters such as the maximization of the overall heat transfer coefficient and minimization of the total pressure drop. Five geometric design variables which include the tube inside diameter, tube outside diameter, pitch size, baffle spacing, and the tube length are investigated for optimization. The governing equations for design and optimization of the shell-and-tube heat exchanger are evaluated and the optimum design parameters are obtained by bees algorithm. The selection of the important design parameters to achieve the proper design is evaluated by fixing each of these parameters, while the other the design parameters are selected as variable to optimize the effectiveness. Compared with the original shell-and-tube heat exchanger, the overall heat transfer coefficient is increased by 22.78% with the minimum increase in the total pressure drop by 1.8%.

A well understanding on the evaporation characteristics and efficient working area of MSHPTRV (multi-stage high pressure and temperature reducing valve) is important for improving the performance and safety of MSHPTRV. The water spraying and evaporation model are integrated into the flow model of MSHPTRV. Compared with the experimental data, the model can show the thermal process well. The flow characteristics and interaction between steam and droplets are presented. On this basis, the increase rate of entropy Sdis is adopted to analyze the thermodynamic loss and innovatively determine the efficient working area of MSHPTRV. The results show that the pressure reducing effect of the second orifice plate is prominent, which accounts for 41.6% of the total pressure drop. The "steam-water layer" is formed at the boundary of steam and water. At the inlet of second orifice plate, the maximum Sdis is 0.782, and the downstream of second orifice plate is the efficient working area of MSHPTRV. The length of evaporation section increases with the droplets diameter significantly.

The spatial structure of a swirling turbulent flow has been investigated based on the three-component laser Doppler anemometry method in an isothermal laboratory model of a four-vortex furnace. The structure of the vortex cores of the flow with the shape of a deformed vertical elliptical cylinder is visualized using the “minimum total pressure” criterion. The spectrum of velocity pulsations indicates the absence of unsteady periodic vortex structures, which means the occurrence of a stable vortex flow in the volume of the combustion chamber.

The gas-particle two-phase flow in a convergent-divergent nozzle is encountered in many applications like solid-propellant rocket motors (SRMs). The “Particle-free Zone” adjacent to the nozzle wall is crucial for the nozzle performance which is closely related to the specific impulse of the motor. Though it has been usually observed by many investigators, the existence of the particle-free zone in different nozzles and its effect on nozzle performance are still ambiguous. In this research, we investigated the two-phase flow with monodispersed micron particles in a nozzle using a two-way coupled Eulerian-Lagrangian model and found that the particle-free zone does not exist for smallest particles (dp = 1.0 μm). Moreover, we built a theoretical model for predicting the extent of the particle-free zone by calculating the trajectory of the particle closest to the nozzle wall. The effects of the particle size, total pressure, total temperature and nozzle geometry on the particle-free zone were studied. Finally, we analyzed the nozzle performance and found that the gas velocity at the outlet and the thrust do not change with particle size monotonously, and the medium-sized particles cause the largest thrust loss.

In this paper, the kerosene/air rotating detonation engines(RDE) are numerically investigated, and the emphasis is laid on the effects of total pressures and equivalence ratios on the operation characteristics of RDE including the initiation, instabilities, and propulsive performance. A hybrid MPI + OpenMP parallel computing model is applied and it is proved to be able to obtain a more effective parallel performance on high performance computing(HPC) systems. A series of cases with the total pressure of 1 MPa, 1.5 MPa, 2 MPa, and the equivalence ratio of 0.9, 1, 1.4 are simulated. On one hand, the total pressure shows a significant impact on the instabilities of rotating detonation waves. The instability phenomenon is observed in cases with low total pressure (1 MPa) and weakened with the increase of the total pressure. The total pressure has a small impact on the detonation wave velocity and the specific impulse. On the other hand, the equivalence ratio shows a negligible influence on the instabilities, while it affects the ignition process and accounts for the detonation velocity deficit. It is more difficult to initiate rotating detonation waves directly in the lean fuel operation condition. Little difference was observed in the thrust with different equivalence ratios of 0.9, 1, and 1.4. The highest specific impulse was obtained in the lean fuel cases, which is around 2700 s. The findings could provide insights into the understanding of the operation characteristics of kerosene/air RDE.

In the current state-of-the-art, high-loss flow in the endwall significantly influences compressor performance. Therefore, the control of endwall corner separation in compressor blade rows is important to consider. Based on the previous research of the Blended Blade and EndWall (BBEW) technique, which can significantly reduce corner separation, in combination with a non-axisymmetric endwall, the full-BBEW technique is proposed in this study to further reduce the separation in endwall region. The principle of the unchanged axial passage area is considered to derive the geometric method for this technique. Three models are further classified based on different geometric characteristics of this technique: the BBEW model, Inclining-Only EndWall (IOEW) model, and full-BBEW model. The most effective design of each model is then found by performing several optimizations at the design point and related numerical investigations over the entire operational conditions. Compared with the prototype, the total pressure loss coefficient decreases by 7%–9% in the optimized full-BBEW at the design point. Moreover, the aerodynamic blockage coefficient over the entire operational range decreases more than the other models, which shows its positive effect for diffusion. This approach has a larger decrease at negative incidence angles where the intersection of the boundary layer plays an important role in corner separation. The analysis shows that the blended blade profile enlarges the dihedral angle and creates a span-wise pressure gradient to move low momentum fluid towards the mainstream. Furthermore, the inclining hub geometry accelerates the accumulated flow in the corner downstream by increasing the pressure gradient. Overall, though losses in the mainstream grow, especially for large incidences, the full-BBEW technique effectively reduces the separation in corners.

This study investigates the combustion performance of a number of novel, crescent shaped cavities placed directly in front of the fuel injection point in a scramjet combustor. Cavities are commonly used for flameholding in scramjet combustors, however they are almost exclusively placed downstream of the fuel injection point and their effect on mixing performance is limited. The crescent cavities in this study were previously found to enhance mixing in chemically frozen flow and the purpose of this study is to investigate whether this performance enhancement translates to chemically reacting flow. The performance of the cavities was numerically studied using unsteady Reynolds-Averaged Navier–Stokes (URANS) modelling with hydrogen as the fuel and an isothermal 1800 K combustor wall, and performance was compared to a no-cavity baseline. The cavities were found to enhance combustion by up to 114% and heat release by up to 143% compared to the baseline when hybrid fuelling is employed and by 47.4% and 54.7%, respectively, when no hybrid fuelling is used. Total pressure loss in the cavity cases was similar to the baseline case and viscous drag was lower in the cavity cases due to the presence of combustion and fuel near the wall. Flame penetration was lower in the cavity cases than in the baseline however as a result of the increased fuel concentration on the wall, while mean heat flux is lower in the cavity cases than in the baseline. Ignition is also found to occur more rapidly in the cavity cases and the flow inside the cavity is oscillatory in nature, as was the case in chemically frozen flow. Given that the crescent cavities in this study significantly enhance combustion performance with limited drawbacks, especially when hybrid fuelling is used, they are an attractive option for operational scramjet combustor design.

The blowout characteristics of partially premixed flame during the condition transition in a cavity-stabilized scramjet combustor are investigated by the experimental method. More attention is placed on the unsteady evolution of the turbulent jet flame, as well as its dynamic response with the initial reactive flow states. The Mach number of flow at the entrance of the combustor is 2.52, the total pressure and the stagnation temperature of which are 1.6 MPa and 1486 K. The cavity-stabilized jet flame is blown out once the injection pressure drop reaches a critical value. During the flame blowout process, the most strongly reactive region gradually transits from vicinity of the cavity shear layer to downstream. Accompanied the sharply decrease of combustion intensity, the light intensity drops rapidly until the flame is completely extinguished. The experimental results indicate that the dynamic process of flame blowout depends on the initial reactive flowfields. The jet flame lasts longer and more intermediate products with high temperature reside within the combustor, which is beneficial to ignite the combustible fuel mixed in the air. As the increase of initial combustion intensity, the flame can be maintained in the combustor at a lower limit of global equivalence ratio after the condition transition. The results on the flame blowout limits are of great significance for the design of the condition transition in a scramjet.

In a supersonic intake, the flow decelerates from supersonic to subsonic speed occurs inside the isolator by a series of bifurcated compression waves and followed by an adverse pressure gradient region. This wave phenomenon is referred to as pseudo-shock wave (PSW). The length of the PSW is one of the main consents in designing the Isolator. The flow characteristic of PSW is depending on the several parameters which makes the difficulties in designing the Isolator for each particular engine. Also, the pseudo shock wave strongly affects the performance of the flow device. In the present work, the characterization of the pseudo-shock wave is analyzed using computational fluid dynamics method in a constant area duct. The effect of back pressure on the shock structure and the influence of geometrical cross-sectional area are studied. The results show the nature of pseudo-shock wave is strongly influenced by the blockage ratio. The large aspect ratio duct shows relatively less total pressure loss. An improved correlation is obtained between the pressure ratio and the length of shock train region in terms of boundary layer blockage ratio. The estimated length of the shock train is compared with present numerical and reference experimental results. The results show a good agreement on predicting the length of shock train.

Based on the migratory phenomenon of the puffer and the cone-shaped structures on its skin, the effects of spinal height and tilt angle on the drag reduction characteristics is presented by numerical simulation in this paper. The results show that the trend of total drag reduction efficiency changes from slow growth to a remarkable decline, while the viscous drag reduction efficiency changes from an obvious increase to steady growth. The total and viscous drag reduction efficiencies are 19.5% and 31.8%, respectively. In addition, with the increase in tilt angle, the total drag reduction efficiency decreases gradually; the viscous drag reduction efficiency first increases and then decreases, finally tending to be stable; and the total and viscous drag reduction efficiency reaches 20.7% and 26.7%, respectively. The flow field results indicate that the pressure drag mainly originates at the front row of the spines and that the total pressure drag can be effectively controlled by reducing the former pressure drag. With the increase in low-speed fluid and the reduction in the near-wall fluid velocity gradient, the viscous drag can be weakened. Nevertheless, the drag reduction effect is achieved only when the decrement of viscous drag is greater than the increment of pressure drag. This work can serve as a theoretical basis for optimizing the structure and distribution parameters of spines on bionic non-smooth surfaces.

Increasing cavitation in a flowing liquid in a pipe system is associated with increasing levels of sound or noise and vibrations. It can lead to large loads if steam bubbles implode close to the pipe wall, which can be harmful to the pipe system. It can lead to fatigue due to vibrations and even worse, erosion of the pipe wall. In the worst case, it can lead to a pipe break. Therefore, cavitation is essential to be predicted and prevented in system design. The aim here is to experimentally verify a concept for the elimination of strong cavitation and harmful cavitation induced vibrations. It has been experimentally demonstrated that replacing a conventional orifice plate with three stages of multi-hole orifice plates in series with only a total streamwise distance of two pipe diameters, giving the same total pressure drop, is sufficient for the elimination of strong cavitation. The cavitation induced vibration levels are reduced by almost 500%. The streamwise distance is almost 500% shorter if multiple conventional orifice plates were to be employed. The same conclusions hold for tests performed with β = 0.30 and 0.40. This compact design was successfully applied at a nuclear power plant, and the reduction of the cavitation induced vibration levels was confirmed. The main contribution is the quantification of the vibration levels, how they depend on the type of orifice plates (conventional, multi-hole, and multi-stage-multi-hole), and how they scale with the dynamic pressure. This will be used for the validation of future CFD simulations including Fluid Structure Interaction (FSI) and advanced two-phase models.

The compressor is a critical component that determines the aerodynamic stability of an aero-engine. Total pressure inlet distortion decreases the thrust and shrinks the stability margin, thus inducing severe performance degradation or even flameout. Generally, tip air injection is used to reduce the adverse influence of total pressure inlet distortion on the aerodynamic stability. In the present work, an experimental investigation on the effects of tip air injection on the stability of a two-stage low-speed axial compressor with total pressure inlet distortion was carried out. A flat baffle generated the total pressure distortion at the inlet of the compressor. The stall margin of the compressor was reduced significantly by the total pressure distortion. When the dimensionless insertion depth of the flat baffle was 0.45, the stall margin decreased to 11.4%. Under the total pressure inlet distortion, tip air injection effectively improved the distortion resistance capability of the compressor. The circumferential layout of the nozzle played a critical role in the stability expansion effect of tip air injection under the inlet flow condition of the total pressure distortion. The modal wave disturbance was likely to occur in the distortion-affected region (the low-pressure region and the mixing region). Tip air injection did not inhibit the generation of the modal wave but restrained the development of the modal wave into the stall cell. It improved the low-speed compressor’s tolerance to the modal wave and allowed a higher amplitude modal wave to occur.

The effect of compressor exit swirl angle (θsw) at the intake of an aero engine combustor on the exit temperature non-uniformity (pattern factor) and combustor total pressure loss is investigated. Experiments are conducted in the engine test rig, measuring the gas temperature and pressure at the inlet and exit planes of the combustor. These parameters are measured at distinct locations along the circumferential and radial directions in the engine test facility. Simulations are carried out using RANS based turbulence modeling and reacting flow approach with Ansys CFX commercial code. The predicted results are validated with experimental data at 5° swirl angle. The swirl angle at the combustor intake is further varied from 0° to 15° and 4 cases (0°, 5°, 10° and 15°) has been considered to predict the effect on the combustor pattern factor and pressure loss. The changes in the flow structure inside the combustion chamber for all these 4 cases are reported in detail. The pattern factor varies from 0.34 to 0.49 as swirl angle changes from 0° to 15°. The lowest pattern factor of 0.34 occurs at 10° swirl angle. However a linear increase in combustor total pressure loss from 5.85% to 6.53% is predicted with the change in swirl angle from 0° to 15°.

To estimate the vitiation effects of test air on flameholding and mode transition in a direct-connected dual mode combustor, an experimental study was conducted with two types of air heater (electric heater and methane combustion heater) at various kerosene equivalence ratios. Total temperature, total pressure and Mach number of combustor entrance were matched in comparative tests. The lean blowout limits, wall pressures and mode transitions were compared. For clean air heater, the lean blowout limit occurred at the equivalence ratio of 0.105, while for methane combustion heater, it occurred at 0.135. Considering mode transition, the combustor operated in subsonic combustion mode at kerosene equivalence ratio above 0.30 with clean air; while for methane combustion heater, mode transition from supersonic to subsonic combustion occurred at equivalence ratio of 0.33. This indicated that the test data should be fully evaluated when using a combustion heater, especially in lean blowout limit and mode transition.

The distribution and evaporation characteristics of a transverse vaporized kerosene jet in supersonic flow were experimentally investigated in a ground direct-connected experimental system. Smear-free and high spatiotemporal resolution images of the jet plume were obtained using the pulsed laser sheet imaging method. The incoming flow had a total pressure of 1.0 MPa, a total temperature of 980 K, and an inlet Mach number of 2.0. The boundary and length of the jet plume were identified using the maximum interclass variance method (Otsu). The experimental results show that the kerosene evaporation process is closely related to the fuel temperature and local flow velocity. The two factors competitively influence the phase distribution in the jet plume, resulting in different evaporation distances in the direction perpendicular to the wall. Under subcritical conditions, the jet penetration depths remain basically constant, while the surface wave structures are distinct on the windward side. The jet of critical-state kerosene exhibits a quasi-gaseous flow structure in the near field and undergoes a complex phase transition of secondary condensation and secondary evaporation. According to the different phase states and evaporation degrees, the jet plume is divided into three evaporation regions, providing a complete description of the evaporation characteristics of kerosene at different temperatures in supersonic flow.

The present study is to explore an active mass flow control technology for an emerging bypass dual throat nozzle. A new arc-shaped bypass system has been applied to replace the previous V-shaped bypass for the bypass dual throat nozzle, which can decrease the total pressure loss significantly. The bypass passage has a contraction part with a variable area ratio. The vector control effectiveness is discussed with different contraction area ratios of the bypass passage. The obtained results reveal that friction choking and geometry choking play crucial roles to affect the bypass mass flow, respectively. When the bypass passage is fully open, the choking location of the bypass flow occurs at the bypass exit, owing to the viscous boundary layer along the constant area passage. A contraction area ratio less than 0.5 does not change the choking position of the bypass flow, because the viscous boundary layer is larger than the geometry contraction height. While the contraction area ratio reaches or exceeds 0.5, the geometry choking takes place in the contraction section. The vectoring angle and bypass mass flow ratio decrease with an increase in the contraction area ratio, whereas thrust coefficient, thrust efficiency, and total pressure loss increase.

Shock loss is the primary source of total pressure loss of transonic axial compressors. Reducing the shock by redesigning the geometry of rotor is of great interest for turbomachinery designers. However, the complex flow field involving shock waves, shock-boundary interaction, intense secondary flows, etc., in the compressor makes the design of rotor difficult. The conventional method of design and optimization is computationally intensive and time-costly. This study introduces an inverse design method to design rotor blades corresponding to prescribed isentropic Mach number distributions with no modification of flow-governing equations. An artificial neural network is trained to predict the isentropic Mach number distributions of any deformed blades. Then, with the pattern search optimization, the blade corresponding to the prescribed isentropic Mach number distributions can be achieved. When the aerodynamic parameter database is calculated and the neural network is obtained, this method can design large numbers of blades of changed isentropic Mach number distributions immediately, without modifying the computational fluid dynamics (CFD) flow solver. The design process is fully automatic and efficient. In this study, NASA Rotor 37 is redesigned and optimized as test cases. Some analysis on the influence of blade shape on aerodynamic characteristics of the rotor is represented in this study.

The precooled turbine engine is applied to overcome the limitation of Mach number due to high temperature inlet air. This paper aims to investigate the effect of water injection cooling on the high-temperature intake air. Then, the theory evaluation and Eulerian-Lagrangian multiphase flow method are conducted to explore the thermodynamic process and resistance characteristics of the pre-cooling section built-in even spray apparatus with a drag reduction. Results show that larger amount of the injection flow rate at higher Mach number will deteriorate total pressure loss and flow field uniformity. Evaporation cooling can decrease flow loss by 9.4%–60.7%. Within 27 ms, total-temperature drop is in 14–144 K range with a low total-pressure drop coefficient of 0.56%–1.29%. Especially, mass flow will increase by 1.15%–18.50%. Thus, water injection cooling is conducive to a higher acceleration, as well as for improving the thermodynamics characteristics of inlet air for a turbine engine at a high Mach number.

This paper presents a satisfactory numerical strategy to reliably evaluate the three-dimensional large-scale flow feature of multistage axial compressors in response to complex swirl distortion with acceptable computational cost. Under the theoretical framework of the body force method, the guide vanes of a swirl distortion generator and the multiple blade rows of a two-stage low-speed axial compressor are described by distributed source terms instead of a complex body-fitted grid approach. The key flow structure of the paired swirl generated by the swirl generator and the main distributions of flow angle at the rotor outlet of the first stage captured by the model agree well with experimental results, demonstrating the effectiveness of the numerical strategy. Additionally, the interaction process between the steady-state paired swirl and the compressor is clearly revealed by the study. The intensity of the swirl distortion can be greatly reduced after passing through the axial compressor. However, the swirl has a significant impact on the local blade loading of the first stage, which induces the mass flux nonuniformity as well as total pressure and total temperature distortion. The combined total pressure and total temperature distortion is significantly attenuated near tip and slightly enhanced near hub as it moves through the second stage.

The topology optimization (TopOpt) is adopted in this article to design tip geometries of a turbine blade. Its principle is to modify the porosity, which is one for fluid and is zero for solid. A CFD method coupled with source terms is implemented for numerical simulations. Moreover, to update the porosity, a quasi-sensitivity and an adjoint sensitivity analysis schemes are integrated with the method of moving asymptotes algorithm, respectively. The TopOpt is conducted to design the tip profile of a 2D tip model and of nine axial sections in a 3D turbine blade. Optimized tips are reconstructed and their performance is verified by experiments and CFD. Results indicate that the re-designed novel tip, which has a convex-concave-convex form and is different from flat and cavity tips, has good performance due to getting rid of flow separations above the tip. Hence, the optimized tip can lower total pressure loss in the gap region.

In a radial bed, the uniformity of the pressure drop distribution is investigated by Euler single-phase flow and porous media models under different operating mode (CF-U/Z, CP-U/Z), gas flow rate (120–240 m3/h), particle diameter ((0.5–3)exp-3 m) and bed voidage (0.3–0.6). According to the nonuniform index η, the uniformity relates to these parameters and improves with increasing total pressure drop of particle bed Δps (sum of the pressure drops of particle bed and gas perforation) or decreasing main channel pressure drop Δpg. Comparing the flow fields with/without particles, Δps is approximately equal to the pressure drop of the particle bed with high-porosity Johnson net, which is well calculated by the Ergun equation. Δpg can be calculated by the modified momentum equation containing k. After changing the wall shear stress and gas-solid axial resistance, it is found that the internal generation factors for k include the influence of gas perforation on boundary layer and the existence of gas axial velocity after perforation. Besides, the global/local k hardly changes with the investigated parameters. The local k is a function of axial position or velocity ratio, which changes obviously at the end of the main channel for the existence of a gas stagnation zone.

Axial turbine is a key device which influences the power output and efficiency of the Compressed Air Energy Storage (CAES) system notably. In present study, the effect of chamber roughness on the efficiency and flow loss characteristics of an axial turbine in the CAES system is investigated numerically to find a way to improve the performance and reduce the manufacturing cost. The results illustrate that the efficiency is reduced nonlinearly with the increase of chamber roughness. An efficiency reduction of 0.45% is obtained when chamber roughness is less than 10 μm at design point. However, the efficiency reduction of 1.9% is obtained by the chamber roughness of 50 μm when the expansion ratio is less than 1.9. Internal flow analysis reveals that the passage vortex near upper end wall in the left-part of stator ring is strengthen with the increase of chamber roughness. An efficiency recovery of 0.61% is achieved by “Local smoothing wall” which has wall roughness of 1.6 μm at only 11.26% of the wall area. The isentropic efficiency is also increased more than 0.6% compared to the full smooth wall wall when total pressure ratio is 1.4.

Maldistribution of fluid among parallel channels is one of the main issues in applications of plate heat exchangers. This paper presents an experimental and numerical investigation of the single-phase flow distribution in the brazed plate heat exchangers, but the results can be applied to plate-and-frame and plate-and-shell designs. In the experiments, the pressure profile in the heat exchanger is measured by the probes inserted into the headers. The flow distribution is determined by the measured pressure drop across the channels and the developed in-channel friction factor correlation. The experimental results indicate that in a U-type brazed plate heat exchanger, the channel flow rate first increases for the first several channels near the heat exchanger entrance due to the sudden expansion of flow in the inlet header. For the rest channels, the flow rate decreases with the distance away from the entrance/exit of the heat exchanger. Such a distribution profile is associated with the axial momentum transfer in the inlet header. The influence of the total flow rate on the distribution profile is trivial, but the maldistribution is more severe with an increased number of plates. Two distribution models presented are developed based on the principle of equal total pressure drop for all flow paths. Two models calculate the pressure profile in the headers by 3-D CFD modeling and 1-D mass and momentum conservation equation. The experimental results validate the models.

As an effective method to influence end wall flow field, non-axisymmetric profiled end wall can improve the aerodynamic performance of compressor cascades. For a highly loaded low pressure compressor cascade, called V103, the study found the optimal non-axisymmetric profiled end wall decreases total pressure loss coefficient by 4.57%, 5.48% and 3.04% under incidences of −3°, 0°, and 3°, respectively, compared with those of the planar end wall. The optimal non-axisymmetric profiled end wall changes the structure of secondary flow in hub region, generating a corner vortex near suction surface, inhibiting the development of the passage vortex towards suction surface and reducing flow separation. When the inlet Mach numbers are 0.62 and 0.72, the total pressure loss coefficient decreases by 3.19% and 4.58% for optimal non-axisymmetric profiled end wall compared with those of the planar end wall. Though optimal non-axisymmetric profiled end wall increases total pressure loss near hub region in blade passage under different inlet Mach numbers, the peak value and region of high loss coefficient above 10% span in blade passage significantly decrease. In addition, different incidences affect the secondary flow streamlines and vortex structure near the cascade hub region, however, different inlet Mach numbers hardly change the secondary flow streamlines and vortex structure. In short, the optimal non-axisymmetric profiled end wall shows better aerodynamic performance than the planar end wall for the highly loaded compressor cascade in multi-conditions.

Some rows of steam and gas turbines, with the exception of the first vanes, operate at the inlet flow conditions created by the upstream turbine stages. In particular, the conditions of the flow at the inlet of a low-pressure turbine vanes (LPT vanes) depend on the upstream high-pressure turbine (HPT). The first stages of the gas turbines are, as a rule, cooled and transonic stages with high power losses; therefore, the nonuniformity of the radial distribution of the angles of flow, the total pressure, and the total temperature may be extremely substantial at the LPT inlet. This leads to increased levels of profile and, especially, secondary losses in the LPT rows. Normally, the LPT vanes have a meridional opening along the inner and outer outlines of the flow path that causes the intensification of the secondary flows and, accordingly, an increase in the secondary losses. To minimize adverse effects on the operation of LPT vanes, such as a nonuniform distribution of the flow parameters at the vane inlet, the meridional opening, and the unfavorable aspect ratio, the width of the blade was increased to decrease the meridional angle of the opening and the three-dimensional design of the vanes was applied to reduce the secondary losses. In particular, variants of the LPT vanes of different spatial configurations were investigated, viz., “reversed bow” vanes, bowed vanes, and leaned vanes in a circumferential direction. The radical change in the vane shape was restricted by the design of the cooling system of the blades that comprises a deflector in the inner hollow of the blade designed to enhance the efficiency of the cooling. As a result, a compromise between the aerodynamic characteristics and the restrictions imposed by the cooling system was found.

Rocket-ejector system as a pressurization unit is of critical importance to the rocket-based combined-cycle engine under the low flight Mach number conditions. Due to the restriction of engine scale, an insufficient mixing process exists in the limited mixer length, leading to a non-uniform stream at the mixer exit. This study concentrated on hot-fire tests to demonstrate the flow characteristics of the rocket-ejector system with the limited mixer length. The results show that shortening the length reduces the critical backpressure of the rocket-ejector system. The rocket-ejector system with a mixer length of 6 length-to-diameter ratio to the straight section can resists the critical pressure ratio of 1.25 in which the entrainment coefficient is 3.432, the total pressure ratio is 36.4 and the occupied area ratio of the rocket exit to the straight section is about 32%. Although the non-uniformity of mixed stream at the mixer exit section would reduce the critical pressure to some extent, the limited mixer length also meet the operation requirements of the rocket-based combined-cycle engine with the diffusion and afterburning cycle.

With the increasing attention to small hydropower, research on improving the hydraulic performance of the pump as turbine (PAT) has become more important. Therefore, to study the influence of clocking effects on the hydraulic performance of PAT, numerical simulation and experimental verification of different guide vane clocking positions are carried out in this paper. The angle between the trailing edge of the guide vane and the volute tongue, θ, is set in eight schemes from 0 to 70 deg. The maximum difference in the efficiency of the eight schemes can reach 2.45% at 80 m3/h. When the flow rate is more than 48 m3/h, the scheme with θ of 50° shows the highest efficiency. The regression equations of entropy generation and total pressure loss are established to obtain visual energy loss. The results indicate that energy loss of the stationary components changes significantly under different schemes, which is the main reason for the change in the efficiency of PAT. Besides, the stator-rotor interaction makes the energy loss near the wake region of the impeller change periodically. This research not only reveals the cause of the clocking effect but also provides a new perspective for the optimization of PAT structures.

Effects of a jet emanating from the nose of a generic rocket in the opposite direction of the incoming freestream have been studied to determine the effects of the jet in reducing the drag and the heat loads on the rocket. The effects are comparatively evaluated for different jet exit Mach numbers from ideally expanded nozzles, designed using the method of characteristics, operating at a total temperature of 500 K and at total pressures with predetermined ratios of the freestream total pressure. Axisymmetric computations were conducted with the generic rocket flying at a Mach number of 6.0 and 40,000 ft altitude with zero-degree angle of attack. Results indicate that a jet generated by the Mach 1.5 nozzle, operating at a total pressure corresponding to 1/4th of the freestream total pressure, resulted in the best combined overall drag and temperature reduction. The area-averaged surface temperature and drag were reduced by 52%, and 22%, respectively, in comparison to the body without the flow control.

The current investigation aimed to identify pressure gradients and to study the fully developed flow patterns of oil-gas as a blend in a pipe of internal diameter 50 mm and 6 m length with different orientations of 0, 30, and 45-degree. The study was performed at constant values of liquid superficial velocities 0.052, 0.157, 0.262, 0.314, 0.419, and 0.524 m/s, and inlet superficial velocities of gas were ranged from 0.05 to 4.7 m/s at atmospheric pressure. Two pressure transducers located up and downstream were used to measure pressure drops inside the tested pipe. Flow patterns were derived by using the correlation between pressure gradients and time series, the Probability Density Function of differential pressures, pressure gradients with gas superficial velocities, and total pressure losses with mean void fractions. The flow patterns of oil-gas were observed as a uniform stratified flow in the pipe on a 0-degree orientation at various superficial velocities. Stratified, wavy, and slug flow patterns were observed at 30-degree orientation, whereas, bubbly, slug, and churn flow patterns were observed in the pipe of 45-degree orientation. The experiment also showed that pressure drop gradients decreased with increased void fractions, gas superficial velocities, and degree rotations of the flow lines. Finally, the validation of using pressure transducers as a technique for estimating the flow patterns of two-phase flow showed acceptable results with some kind of patterns.

The increasing performance of modern aeroengines led the research towards the optimization of machine components not deeply analyzed in the past. In this context, the mechanisms driving the interaction process between the secondary flows evolving at the hub of low-pressure turbines with the rotor-stator cavity systems have been poorly investigated in the literature. In this work, an experimental and numerical analysis of the interaction between the endwall near wall flow and the leakage flow of a real cavity system is presented. The experimental results were carried out in the annular low-pressure axial flow turbine of the University of Genova. Experimental blade loading and pressure distributions into the cavity, as well as the measured total pressure loss coefficient, have been used for a proper validation of CFD results. Both steady and unsteady calculations were carried out through the commercial solver Numeca. Particularly, several numerical approaches have been tested into this work: RANS, Non Linear Harmonic (NLH), and URANS. The most promising CFD techniques have been firstly identified by comparison with experimental results and then systematically employed to extend the analysis of secondary flow-cavity flow interaction to positions and quantities not available from the experiments. Losses characterizing the mean flow-cavity flow interaction process will be shown to cover a great amount of the overall stage losses and should be properly accounted for the design of future optimized cavity configurations.

Parabolic flight experiments were carried out to investigate the effect of ambient pressure on the extinction limit for opposed flame spread over an electric wire insulation in microgravity. Low-density polyethylene insulated Nickel-Chrome wires with inner core diameter of 0.50 mm and insulation thickness of 0.30 mm were examined for ambient pressures ranging from 50 kPa to 140 kPa for an external opposed flow of 10 cm/s. The experiments showed that the limiting volumetric oxygen concentration (LOC) increased as the total ambient pressure decreased. This LOC trend can be explained by radiation loss from wire surface. The radiation loss increased as the ambient pressure decreased – a result that can be explained by the increase in preheat length with decreasing ambient pressure. Moreover, when the data was plotted in a partial pressure vs. total pressure space, it became evident that the limiting oxygen partial pressure (LOPP) decreased with decreasing total ambient pressure. This LOPP trend can be explained by the fact that the flame temperature increased as the ambient pressure decreased under constant oxygen partial pressure. In current fire safety design for spacecraft, tentative oxygen concentration criteria in spacecraft suggested by NASA is assumed as 30% of oxygen concentration, and this value is assumed independent of ambient pressure. However, the present result implies that consideration of the effect of ambient pressure on the flammability limit is necessary, especially with respect to the possibility of an extension of the allowable atmosphere condition for spacecraft cabin in the low pressure region.

The mixing loss in S-shaped duct causes the contradiction between the total pressure recovery and circumferential distortion, which confuses the design process for boundary layer ingestion (BLI) inlet. This article is aiming to present a detailed understanding of the effect of two kinds of loss resources (mixing loss and frictional loss) on the performance for various BLI inlet geometries. To achieve the above objective, a design method including three control parameters is established. With this method, a systematic investigation of control parameters on the performance of BLI inlet is carried out, which leads to the contour map for the performance of BLI inlet. Based on the map, the flow physics of BLI inlet with two kinds of losses is then analyzed. Finally, an empirical design principle is provided. The results show that the appropriate design for BLI inlet has the potential to increase the total pressure recovery or reduce the circumferential distortion.

Increasing the fuel mixing performance in the combustor of scramjet substantially improves the overall efficiency of the scramjet engine. In this article, computational fluid dynamic is used to study the impacts of hydrogen jets injection through the backward-facing multi-steps on the fuel distribution and mixing zone at the supersonic air stream of Mach = 4. This study also analyzes the jet flow feature and circulation of jets in different sections of the combustor at downstream of the multi-injectors. Reynolds average Navier-Stocks equations are solved with SST turbulence model for achieving a precise and acceptable solution. The effects of step height on the jet features are also examined. According to circulation evaluation, low jet total pressure (pressure ratio = 0.1) and high step depth (step depth = 1 mm) is the optimum condition for achieving high circulation value. Our investigations show that the mixing efficiency of the hydrogen multijets improves up to 15% when the step height increases from 0.5 mm to 1 mm.

The windage loss caused by protrusion in a rotor–stator cavity has been studied in detail, and there are abundant fitting formulas that have been summarized to calculate the moment coefficients. Some other theorists have emphasized its effect on the sealing efficiency, proposing that installation of protrusion could alleviate gas ingestion. However, the protrusion shape which is an influential factor on the sealing efficiency has not been focused in previous research. Using the experimental method of measuring CO2 volume fraction, cavity pressure, and power consumption, we investigated the effects of several typical protrusion shapes on various parameters for two sealing structures, in order to obtain the optimal shape. Results showed that a variation of the protrusion shape had little impact on the static pressure, but the total pressure and the sealing efficiency increased in different degrees. Furthermore, even though the hexagon shape resulted in the highest sealing efficiency, we observed that the drop shape had the best overall performance in all of the eight models, which could result in higher efficiency of the turbine cavity. The combination of a radial seal structure and protrusion could improve sealing efficiency better.

In this work is presented a hydrodynamic and mass transfer study using the computational fluid dynamic (CFD) for typical experimental measurements and theoretical calculations for pyrite (FeS2) dissolution in acid media using ozone as an oxidizing agent, which occurs in a laboratory glass reactor. The hydrodynamics for the reactor employed, was evaluated with the Navier-Stokes equations (κ-ɛ) RNG turbulence model and the VOF multi-phase model. The results of the hydrodynamic simulation study show that the reactor can operate at a speed range of 1.2 to 4.4 m/s at a total pressure of 12.5 Pa; the Reynolds numerical profile indicates that the flow has a turbulent behavior inside the reactor. For the mass transfer simulation, the model of finite velocity-Eddy dissipation was used for solving the kinetic of the reaction, and was performed for 3 different particle fractions: -106 +75 µm, -38 +25 µm and -25 µm. The simulation results present satisfactorily differences of 17%, 13% and 2% respectively, in comparison with experimental data for pyrite dissolved.

The use of a submerged inlet is advantageous in modern aircrafts because of its low drag resistance, small radar cross section and ease of maintenance. Although it is well known that the forebody boundary layer deteriorates the aerodynamic performance of a submerged inlet, the level of impact has not yet been fully quantified. To quantify the forebody boundary-layer effect, a submerged diverter was designed to remove a portion of the low-energy boundary flow. The flow pattern and aerodynamic performance of a submerged inlet, with and without the diverter, were investigated by wind-tunnel experimentation and numerical simulations. The effects of mass flow, free stream speed, angle-of-attack and sideslip angle on the aerodynamic characteristics of the inlet with and without the submerged diverter were studied, over an operating envelope of M0 = 0.3 ∼ 0.6, $\alpha$ = –6$^{\circ}$ ∼ 8$^{\circ}$ and $\beta$ = 0$^{\circ}$ ∼ 4$^{\circ}$. The results indicate that both the total pressure loss and the circumferential distortion can be significantly reduced with the removal of the forebody boundary-layer low-energy flow. Meanwhile, the main mechanisms for losses in the submerged inlet were also analysed.

The Authors presented an interesting comparison of 3D Navier–Stokes solvers for dam-break-like flows using two different modules for the pressure, one that calculates a total pressure solving a Poi...

Theoretical and experimental studies were carried out 10 investigate the effects of duct velocity, sin density and tip-to-shroud clearance on the flow bypass and its impact on the pressure drop across a longitudinal aluminum fin array and its thermal performance. The clearance was varied parametrically, stating with the fully shrouded case and variations of the channel height giving partially shrouded configuration of different clearance ratios were also carried out. The slow bypass was found to increase with increasing sin density and insensitive to the air slow rate. That effect of fin density decreased as the clearance increased. The calculated total pressure was greatly affected by fin density. For fully-shrouded sin array, with H/S equals to 8 and 12.72, the pressure drop increased by a factor of 4.3 and 20 of that with Hf/S equals to 3.4, respectively. The total pressure drop and the average convective heat transfer coefficients corresponding to the fully and partially shrouded fin array of Hf /S=3.4 were compared. Going from fully to partially shrouded one of the largest clearance ratio (C/Hf =0.89), the total pressure drop reduced by about 50%. For clearance ratios equal to 0.36, 0.56, and 0.89, the average heal transfer coefficients were reduced by about 12, 17, and 30 percent of those for the fully shrouded configuration at ReD of about 3x103. That percentage reduction in heat transfer coefficients decreased with the increase of air flow rate.

Progressing advancement concerning the intensification of engine thrust to weight ratio in a gas turbine engine offers ascends to extremely loaded compressor stages which have a significant concern in recent times. Non-uniform Inlet flow conditions are frequently occurring in aircraft gas turbines engines that result in chronological distortions of total temperature and total pressure at the engine inlet. Distorted inlet flow operation of the axial compressor deteriorates the aerodynamic performance which reduces stall margin and increased blade stress levels. The current paper aims at the investigation of the flow field in the tip clearance region of the transonic micro-axial flow compressor at non-uniform flow conditions. Under the design and off-design RPM, the current paper focuses on the adverse effects of a steady-state distorted pattern of total pressure inlet flow conditions, including hub radial, tip radial pressure distortion, and combined effects of pressure distortion with bulk swirl flow to evaluate its effects on compressor performance and stability. The results show that like hub radial pressure distortion improves the stability range of turbofan transonic micro-compressor rotor. Whereas tip radial and counter swirl flow deteriorate the performance of the compressor. Furthermore, compressor performance alleviated the stall at higher RPM combined with distorted inlet pressure.

The important operating characteristics of pulsed Pressure Gain Combustion (PGC) propulsion are the pressure gain of the combustor component and the propulsive performance gain of the engine. A ramjet-type valve/valveless air-breathing pulsed detonation engine with a supersonic internal compression inlet is investigated. Based on an ideal thermal cycle, the ideal equivalent pressure ratios (πcb) of the Pulsed Detonation Combustor (PDC) are obtained theoretically which are directly related with the propulsive performance of the engine. By introducing an orifice loss model into the cycles, the critical pressure drop ratios through the orifice for the PDC achieving pressure gain and the engine achieving thrust gain are studied. More influencing factors are investigated by the use of a one-dimensional (1-D) numerical simulation model. The operating characteristics of the pulse detonation engine are investigated with changes of the valve type, the inlet/outlet area ratio of the PDC, the nozzle area ratio, and flight conditions. All these factors can affect πcb of the PDC, and πcb can be optimized by changing the geometry of the engine. The most important influence parameter is the valve type. When using an orifice-type aerodynamic valve, simulation results show that the PDC cannot achieve the pressure gain characteristics. When a supersonic internal compression inlet is introduced to the engine, whether the Pulse Detonation Engine (PDE) can achieve thrust gain comparable with that of an ideal Brayton cycle engine not only is related to the pressure gain of the combustor, but also needs to optimize the engine structure to reduce the total pressure loss.

Detonation combustors are considered promising alternatives to conventional combustors because they offer high thermal efficiency and fast combustion. However, especially for the rotating detonation combustor, the theoretical propulsive performance has not been confirmed in experimental studies because the highly unsteady flow field hinders the measurements process. To understand the involved phenomena in more detail, a reflective shuttling detonation combustor (RSDC) with a rectangular combustion chamber was developed. The interior of the chamber can easily be visualized owing to its two-dimensional quality. Utilizing the RSDC, several combustion tests with gaseous ethylene and oxygen were conducted for different values of mass flow rates and equivalence ratios. Combustion modes from the tests were classified into four types based on the fast Fourier transform (FFT) analysis of the luminous intensity of the CH* self-luminescence images captured by a high-speed camera and a band pass filter. Simultaneously, the theoretical total pressure of a conventional isobaric combustor was compared to the static pressure measured at the bottom of the RSDC chamber. For the detonation modes, the ratio between experimentally measured static pressure and the theoretical pressure varied depending on the location in the chamber owing to the distribution of the time-averaged static pressure. Furthermore, the pressure ratio of the detonation modes was up to 18% lower than that of the deflagration mode potentially owing to the flow velocity induced by the detonation waves.

Achieving combustion in the available residence time by efficient mixing of the injected fuel has been a challenge for scramjet designers since years. Design of the injection configuration do affect the patterns of mixing and in turn the rate of heat release for a given combustor geometry. The presented work attempts to investigate the effect of inclined injection configurations on the performance of a scramjet for a typical combustor geometry having intrusive fuel injection using strut. Over a decades of literature points towards the fact that numerical simulations are widely being accepted as promising method of investigation of flow inside a scramjet. Reacting simulations of hydrogen injection in an inlet air stream of Mach 2.5 are done with various fuel injection configurations. After validation of simulation results with the available experimental Sadatake Tomioka et al. (2001) results, different injection configurations were simulated for a constant equivalence ratio. The maximum combustion, mixing and heating was observed in near the strut surface, hence flow fields of these configurations are studied near the strut to understand the mechanism of mixing. Observations on performance parameters like total pressure loss, maximum top wall static pressure, flow enthalpy etc. are made to compare these configurations.

The combustion instabilities of supersonic combustion were investigated experimentally in a laboratory-scale scramjet combustor with a cavity flame holder. Ethylene was injected transversely from an orifice to the supersonic flow of Mach 2 with a stagnation temperature of 1900 K and a total pressure of 0.37 MPa. The dynamic pressure, CH* chemiluminescence and shadowgraph images were measured with a pressure sensor and a high-speed video camera. Dynamic pressure was analyzed by fast Fourier transform, and time-resolved CH* chemiluminescence images were modally decomposed by the sparsity-promoting dynamic mode decomposition (SP-DMD). The results indicated that two combustion instabilities were observed for cavity shear-layer stabilized combustion and the oscillation between jet-wake stabilized and cavity shear-layer ram combustions for the power spectral density (PSD) of pressure. In the case of the combustion instability of cavity shear-layer stabilized combustion, a dominant peak of approximately 128 Hz was observed for the PSD of pressure. This instability corresponded to an entire flame oscillation of the cavity shear-layer stabilized combustion, which was validated by the SP-DMD and a low rank reproduction with 10 modes. This was driven by a fuel injection oscillation in the injection orifice. In the case of oscillation between the jet-wake stabilized and the cavity shear-layer ram combustions, peaks around 1600 Hz were observed for the PSD of pressure. This mechanism was also explained by the SP-DMD modes and a low rank reproduction using within 10 modes. The DMD and shadowgraph images indicated that the vortex formed by a separation of the boundary layer induced a strong jet-wake flame, resulting in the temporal thermal choke followed by cavity shear-layer stabilized ram combustion. The data-driven approach with SP-DMD clarified the combustion instability mechanisms of the supersonic combustion in detail.

Plasma jet has been widely used in supersonic combustor as an effective ignition and combustion assisted method, but currently it is mostly combined with the traditional wall fuel injection method, while the application combined with the central fuel injection method is less. In order to expand the combustion range, the plasma jet was introduced into a strut-cavity combustor with an alternating-wedge. The effects of total pressure of strut fuel injection, total pressure of cavity fuel injection, total pressure of plasma jet injection and plasma jet media on the combustion characteristics were analyzed in supersonic flow by numerical calculations in a three-dimensional domain. The combustion field structure, wall pressure distribution, combustion efficiency and distribution of H2O at the exit of the combustor with different injection conditions were analyzed. The results show that the combustion efficiency decreases with the increase of the strut fuel injection total pressure. However, the combustion area downstream increases when the total pressure of the strut fuel injection increases within the proper range. The combustion range is expanded and the combustion efficiency is improved when the cavity fuel injection total pressure is increased within the range of 0.5–2.0 MPa, but a sharp drop in combustion efficiency can be found due to limited fuel mixing when the total injection pressure of the cavity fuel is excessively increased. With the increased total injection pressure of the plasma jet, the height of the cavity shear layer is raised and the equivalence ratio of the gas mixture in the cavity is improved. When the total pressure of the plasma jet is 1.25 MPa, the combustion efficiency reaches a maximum of 82.1%. The combustion-assisted effect of different plasma jet media is significantly different. When the medium of the plasma jet is O2, the combustion-assisted effect on the combustor is most significant.