Inflow Mach Number Research Articles

Investigations of shock–boundary layer interaction dynamics using high-bandwidth pressure field imaging

The large-scale pulsations of shock-induced separation with length scale that significantly exceeds the incoming boundary layer thickness are investigated. The shock–boundary layer interaction (SBLI) unit is generated by an inward-turning axisymmetric compression ramp at an inflow Mach number of 2.5. A substantial region surrounding the centre azimuth exhibited mean and dynamic flow features that are consistent with two-dimensional separation. Two-dimensional highly resolved maps of surface pressure field are obtained using fast-response pressure-sensitive paint fluorescence imaging at 40 kHz repetition rate. The measurement domain covered significant regions of the incoming boundary layer through the relaxing boundary layer downstream of the reattachment as well as over 25 boundary layer thicknesses in the azimuthal direction. These measurements provide new insights into the spanwise coupling of the SBLI unit in addition to its inherent dynamics. The power spectral density (PSD) of the centreline pressure exhibits very good agreement with theoretical models and complementary measurements using fast-response pressure transducers, which served to validate the pressure field measurements. Detailed examination of the PSD reveals strong agreement with the literature, which includes the peak Strouhal number of the separation and reattachment shock motions as well as the downward frequency shift along the separation bubble. Furthermore, the pressure fluctuation maps reveal streamwise-elongated structures just downstream of the ramp leading edge that persist well downstream of the reattachment. A time sequence of conditional average pressure fluctuation maps is constructed surrounding isolated pressure excursions in the intermittent region. This sequence, along with two-point cross correlation analysis, provides critical information about the flow processes that drive the separation bubble pulsations in the SBLI units with large separation scales. Overall, the imbalance in the mass within the separation bubble appears to be the critical mechanism that drives the separation bubble pulsations. Furthermore, the pressure perturbations originating at azimuthally offset locations are also observed to influence the separation bubble dynamics.

Theoretical profiles of two-phase transcritical laminar boundary layer near the Widom line

Boundary layers can be defined as sub-, trans-, and supercritical regime according to whether their temperature profiles cross the Widom line. The transcritical boundary layers which contain liquid and supercritical phases are hard to solve. In this paper, adiabatic flat plate laminar boundary layers of supercritical carbon dioxide (SCO2) are preliminarily explored by the self-similar solution. We analyze the difference of boundary layer profiles at 8 MPa - inflow pressure by varying inflow temperature Te and Mach number Ma. The influence of Te and Ma on the velocity profiles exhibits complex diversity at three boundary layer regimes. Compared with subcritical and supercritical boundary layers, transcritical boundary layers with smaller skin friction coefficients are the most special, and they include the “sub-layer”. Notably, there is no transcritical boundary layer when Te increases to pseudo-critical temperature Tpc at small Ma. The wall temperature Tw of transcritical boundary layers may be greater than inflow total temperature T0 at larger Ma.

Steady-state error estimation and calibration of a shielded total temperature probe for turbomachinery

Accurate measurement of the total gas temperature plays a vital role in the performance research of turbomachinery. The total temperature thermocouple (TC) used in the measurement often needs to be specially designed and calibrated. In the past, the calibration was usually completed by a complex and highly costly experiment. The present work developed a new shielded total temperature probe (NTTP) for relatively low-temperature gas measurement. Meantime, this new design's characteristics and different error sources are analyzed based on the Conjugate Heat Transfer (CHT) simulation method. The CHT results show that the thermal conduction error is kept at a small value in the range of Mach number from 0.2 to 0.6; The velocity error contributes more than 90% of the total error. Four standard temperature probes are selected for the comparison. The maximum measurement error is 1.13 K, which is 331.62%, 119.4%, 61.6% and 59.5% lower than the standard type-A, B, C and D, respectively. Moreover, the detailed flow field and the influence of TC junction position on measurement accuracy are discussed. With the increases of the junction away from the supporting, the total measurement error of the NTTP increases first and then decreases. To vindicate the applicability of the proposed design, the measurement error of the NTTP is verified by a new experimental method with the relatively static fluid as the gas flow source. The testing results show that when the inflow's Mach number is 0.52, the measurement error of the NTTP is about 40% lower than that of the type-C total temperature probe.

Unsteady Oblique Detonation Waves in a Tunnel Induced by Inflow Mach Number Variation

Oblique detonation waves (ODWs) have been investigated widely aiming at facilitating their application in hypersonic engines. However, there is a lack of research on unsteady ODWs which are unavoidable in the hypersonic air-breathing scenario. In this study, unsteady ODWs triggered by the variation of the inflow Mach number (M0) have been studied and the geometric model is a tunnel with an outward-deflection upper wall to mimic an engine outlet. Numerical results demonstrate that when M0 deviates from the designed state, two typical wave structures arise, featuring a Mach stem of detonation or a post-corner recirculation zone. A sudden change in M0 leads to the transition of these two structures, generating unsteady ODWs temporally with a multi-segment-complex wave surface caused by triple points. The wave structures near the corner have been analyzed in detail, revealing how the Mach stem and the recirculation zone evolve into each other. Furthermore, the effects of unsteady ODWs on hypersonic propulsion applications have been discussed, providing possible ways to suppress the Mach stem of detonation.

Investigation of Mixing Characteristics of the Transverse Jet in a Supersonic Combustor at Low Mach Inflow Condition

LES is used in the present work to simulate and unfold unsteady flow dynamics and mixing characteristics in a cavity-based supersonic combustor at low Mach numbers. In the combustor, a transverse jet is injected ahead of the cavity. Two inflow Mach numbers (Ma=1.2 and Ma=2.5) are numerically investigated. Fine instantaneous and average mixing flow structures are well identified and analyzed by LES. The mixing performance is also studied and compared. The simulation results show that penetration is dominated by jet/freestream momentum flux ratio, even though the inflow Mach number varies. It is also demonstrated that for a jet at low Mach inflow, a higher coefficient of pressure is attained. While around the jet location, a much higher peak of the mixing vorticity efficiency is obtained for low Mach inflow, which is of fundamental importance to the near-field mixing.

Effect of a streamwise vortex on supersonic boundary layer transition

Supersonic boundary layer transition induced by a streamwise vortex is experimentally investigated based on the nanoparticle-based planar laser scattering technology, with an inflow Mach number of 2.7 and a unit Reynolds number of 8.79 × 106 m−1. Though the vortex is outside the boundary layer, it results in obvious difference of transition position along spanwise direction. On the up-wash side, transition to turbulence is completed in advance compared to the vortex-free case, and transition position moves more upstream as the vortex approaches the wall. Conversely, the transition process on down-wash side is inhibited. The promotion and inhibition effects on transition are found to be induced by the influence of vortex on streak density.

Numerical study of the effect of a sudden change in inflow velocity on the stability of an oblique detonation reflected wave system

<p indent="0mm">The oblique detonation engine (ODE) is expected to be used in future hypersonic power systems, bringing many technical advantages to the engine by regulating the combustion mode. However, few studies concern the wave structure of an oblique detonation wave (ODW) in confined spaces, particularly for the stability of reflected wave systems under complex inflow conditions. On the basis of a simplified ODE combustor model, the effects of a changing inflow velocity on the steadiness of the oblique detonation reflected wave system are studied in this paper. The numerical results show that the evolution of the reflected wave system undergoes two processes, including long-term accumulation and sudden acceleration, which is unfavorable for an ODE application and can be efficaciously harnessed by increasing the inflow Mach number. The inflow disturbance aroused by this increase may restabilize the originally unstable reflected wave system, and the effects depend on the time and amplitude of the increase of the inflow Mach number. The earlier the disturbance occurs and the greater the increase in Mach number, the more easily a stable reflected wave system will develop. Furthermore, by analyzing the flow field characteristics of ODWs disturbed by inflow, the re-stabilization mechanism of a reflected wave system is revealed; that is, the flow choking induced by the subsonic zone behind the Mach stem is suppressed by the inflow disturbance.

The effect of mixing and heat transfer on regression rate of TAGN-based fuel in a segregated AP/TAGN solid motor

The effect of mixing and heat transfer on regression rate of TAGN-based fuel of a segregated AP/TAGN solid motor is studied in the present study. A regression rate model of TAGN-based fuel under coupled combustion is developed and validated based on detailed reaction mechanism. The two-dimensional mixing characteristics in boundary layer and the effect on regression rate of TAGN-based fuel are analyzed in detail. The simulation results show that the increase of initial temperature of oxidizer gas leads to better mixing in the axial direction and higher temperature of burning surface. However, the mass concentration of O2 decreases greatly due to the increasing consumption by primary combustion reactions of AP-derived gaseous species. Therefore, the regression rate of burning surface decreases slightly. With inflow Mach number increasing from 0.32 to 0.75, mixing in the boundary layer is greatly enhanced. Therefore, both the temperature and regression rate increase with inflow Mach number. As pressure increases from 4 MPa to 10 MPa, the primary reactions of AP-derived gas species are also enhanced. Therefore, the mass concentration of O2 decreases rapidly, which leads to oxygen-lean combustion in the combustion chamber. Therefore, the temperature of burning surface decreases while the regression hardly varies despite of the better mixing near the burning surface at high pressure. The results show that the coupled combustion reactions between inflow oxidizer gas and fuel gas mainly occur in a narrow belt-like region of the boundary layer. The results show that the regression rate of solid fuel in the segregated solid motor is controlled by the competition between primary combustion and coupled combustion.

Relationship between physical parameters of supercritical fluids and normal shock characteristics

Physical parameters of supercritical fluids change drastically near the critical region, which makes it difficult to predict and analyze the supercritical fluid flow parameters behind the normal shock wave. In this paper, in combination with supercritical fluid physical parameters database, we employed an iterative algorithm to solve the flow parameters behind normal shock by deriving shock equations. The change of normal shock parameters of six supercritical fluids with inflow state was studied by the controlled variable method and a correlation analysis. The results show that when the inflow Mach number is fixed, the normal shock parameters, such as density ratio and pressure ratio, change rapidly in the Widom zone as a result of the dramatic changes of the physical parameters. When the inflow state is the same, the normal shock pressure ratio of NH3 is the highest, whereas that of C8H24O2Si3 and C10H22 are pretty low. The normal shock intensity of supercritical fluid is better reflected by the pressure ratio rather than Mach number. According to the correlation analysis, the compressibility factor and the sound speed of inflow are the main physical factors that affect the normal shock density ratio and pressure ratio, respectively. Based on the main physical factors, empirical equations for predicting the change trend of normal shock pressure ratio and density ratio are obtained.

Experimental and Numerical Investigations on the Mixing Process of Supercritical Jet Injected into a Supersonic Crossflow

The mixing process and distribution characteristics of a supercritical endothermic hydrocarbon fuel (EHF) jet injected into a supersonic crossflow were investigated by experimental and numerical methods, respectively. The schlieren system and acetone planar laser-induced fluorescence (PLIF) optical system were used to capture the flow-field structural characteristics and instantaneous plume. The mixture and real gas models were employed to calculate the interaction of a transverse jet and supersonic crossflow and reveal a good accuracy with the experimental results. The mixing efficiency and total pressure loss were analyzed based on the numerical results. The results indicate that the supercritical-state EHF directly changes to a gaseous state as it enters the supersonic crossflow from the injector. The EHF jet plume boundary increases with the increasing momentum flux ratio (q). As the streamwise and spanwise distance increases, the traverse heights and expand width increase, and the EHF jet plume presents a semicircle shape in the cross-sectional plane. With the increase in the traverse direction, the concentration distribution shows a fast and then slow power exponential decreasing law; the highest concentration point starts from the near-wall region and rises in the transverse direction with the flow distance increasing. For the same injection condition, the higher the inflow Mach number, the higher the mixing efficiency. For the same Ma, the mixing efficiency is better for the case with low injection pressure and high injection temperature. The total pressure loss is greater in the higher Ma, and high injection pressure conditions cause greater total pressure loss.

Simulation of high-Mach-number inviscid flows using a third-order Runge-Kutta and fifth-order WENO-based finite-difference lattice Boltzmann method

A discrete-velocity Boltzmann equation(DVBE) with Bhatnagar-Gross-Krook (BGK) approximation is discretized in time and space using a third-order Runge-Kutta (RK3) and fifth-order weighted essentially nonoscillatory (WENO) finite-difference scheme to simulate benchmark inviscid compressible flows. The implementation of the WENO ensures that solutions behave with minimal or no oscillations, narrowing the gap between the exact and the numerical results. Discrete-velocity sets given by Kataoka and Tsutahara [Phys. Rev. E 69, 056702 (2004)10.1103/PhysRevE.69.056702] are used. The additional dissipation terms as well as artificial viscosity are incorporated in the formulation to solve the compressible flows at high Mach number. Further, the flows which are subjected initially to a high density ratio are effectively simulated. In this article, one-dimensional benchmarks are simulated at initial Mach number up to 30 and density ratio up to 1000, whereas, the benchmarks in two dimensions are simulated with a Mach number up to 10. The algorithm is assessed by simulating numerous benchmarks, namely, (i) one-dimensional Riemann problem for various shock waves combinations [namely (a) shock-shock waves in the case of different Mach numbers, (b) rarefaction-shock waves for various density ratios, (c) sudden contact shock discontinuity, and (d) shock-rarefaction waves for density ratio 5], (ii) isentropic vortex convection test, (iii) regular shock reflection (RR) for Mach numbers 2.9 and 10, (iv) double Mach reflection (DMR) for inflow Mach numbers as 6 and 10, and (v) forward-facing step for inflow Mach numbers 2 to 5. Further, the effect of change in Mach numbers and wedge angles on the flow structures in the case of DMR are detailed. In the case of a forward-facing step, the variations of flow structure (e.g., the Mach stem height, triple points location, and shock standoff distance) are detailed with respect to Mach number, step height, and specific-heat ratios. Finally, the numerical stability of the proposed formulation is carried out to assess the behavior of the free parameters.

Numerical investigation on movement of triple points on oblique detonation surfaces

A normal detonation wave in a gaseous mixture is a transient, multidimensional structure containing triple points (TPs) that collide in pairs and then propagate oppositely. However, the TPs on an oblique detonation wave (ODW) almost propagate along the same direction in most studies. In this study, the reactive Euler equations coupled with a two-step induction–reaction kinetic model are used to solve a two-dimensional wedge-induced ODW. Two novel movement patterns are observed in most cases. Results show that the TPs of the ODW can propagate upstream and even stand on the wave surface. The movement patterns of TPs include downstream, upstream, and steady according to their propagation direction relative to the wedge. We find that the ratio of the post-ODW flow speed Uτ to the transverse wave speed UT dominates the TP movement types. When the speed ratio Uτ/UT is approximately equal to 1, the TPs can stand on the wave surface. Above unity, downstream TPs form, and upstream TPs correspond to a value smaller than 1. Furthermore, the inflow Mach number has little influence on UT, while Uτ changes significantly. This is largely due to the high sensitivity of the ODW angle to the inflow. The high heat release rate benefits upstream TPs, and steady TPs form under a large wedge angle. The results are confirmed by varying the inflow Mach number, wedge angle, and chemical parameters.

Investigation for Effects of Jet Scale on Flame Stabilization in Scramjet Combustor

Jet scale affects the mixing and combustion of fuel and inflow. With the increase in the scale of scramjet combustors, the study of large-scale jets is particularly significant. The effects of jet scale on flame stability in scramjet combustors were studied by direct-connect combustion experiments. In this paper, the flame distribution characteristics of different jet scales were compared by using similar jet/inflow momentum ratios. The inflow Mach numbers were 2.4 and 3.0, and the total temperature was 1265 K and 1600 K, respectively. The results show that, when the equivalence ratio increases, the combustion intensity increases. Under the condition of same momentum ratio, the increase of jet scale is conducive to fuel injection into the core mainstream, increasing heat release, and the flame stabilization mode will change from cavity stabilization mode to jet-wake stabilization mode. Increasing the distance between jet orifices is not beneficial to combustion, and may even lead to blowoff.

Electromagnetic field/hypersonic flow field coupled algorithm and its application in the magnetic controlled inlet design

A novel magnetic controlled inlet design method is proposed for the airbreathing hypersonic vehicles cruising at an extremely high Mach number. By introducing the electromagnetic source terms to the Euler equations, an electromagnetic field/hypersonic flow field coupled algorithm is developed as the numerical design tool. A quasi-one-dimensional simplified model of the magnetic controlled inlet is constructed to analyze the compression performance under the different combinations of magnetic field and leading-edge shock angle. The results show that the combination of weak leading-edge shock and strong magnetic field can obtain a better compression effect. Three inlets are designed under the magnetic field with the induction intensity of 0.05 T, 0.08 T, and 0.1 T, respectively. The designed inlets have a contraction ratio of nearly 1.5 and the wall shape is not sensitive to the variation of magnetic induction intensity. With an inflow Mach number of 10, the exit Mach number can be reduced from 5 to 2.5 when the magnetic induction intensity is adjusted from 0.05 T to 0.1 T. In addition, there is no reflected shock in the inner channel of the inlet and therefore the complex shock/boundary layer interaction in the conventional inlet has been greatly relieved. Although the total pressure recovery is significantly reduced, part of the energy loss can be transformed to electric energy by the MHD energy extraction equipment and used to accelerate the nozzle flow.

Rarefaction Effect, Heat Transfer, and Drag Coefficient for Gas Flow Around Square Cylinder in Transition Flow Regime

Abstract In this work, rarefaction effect, heat transfer, and drag coefficient for gas flow around a square cylinder in transition flow regime are numerically studied using unified gas kinetic scheme (UGKS). To reduce computational cost, a mirror symmetry boundary treatment for UGKS is proposed and applied in this study. It is found that: (1) velocity slip is not obvious on upwind and downwind surfaces of square cylinder due to the tangential gas flow being weak in these surfaces; (2) with the increase of local Knudsen number, velocity slip on upper surface always increases, but temperature jump can decrease, which indicates that Knudsen number is not the decisive parameter to characterize temperature jump; (3) the heat transfer between gas and square cylinder enhances with the increase of inflow Knudsen number and cylinder temperature due to the increase of temperature jump and deteriorates with the increase of inflow Mach number on account of gas stagnation; and (4) the drag coefficient increases with the increase of inflow Knudsen number, the decrease of inflow Mach number, and the increase of cylinder temperature. To further predict the variation of average Nusselt number and drag coefficient, correlations for average Nusselt number and drag coefficient with inflow Mach number ranging from 0.05 to 0.3, inflow Knudsen number ranging from 0.1 to 10, and cylinder temperature ranging from 320 K to 380 K are proposed. This research can improve the understanding for mechanisms of gas flow and heat transfer in micro-electromechanical system (MEMS) devices.

Influence of uncertain inflow conditions on a subsonic compressor cascade based on wind tunnel experiment

The intended inflow condition of compressors is subject to the uncertain inflow Mach number ( Ma1) and incidence angle ( i). To quantify the effects of uncertain inflow conditions, real-time measuring experiments of Ma1 and i are performed under various operating conditions to extract the statistical characteristics of the uncertainties. Statistical analysis shows that Ma1 and i follow Gaussian distribution. 0.005 and 0.125 are chosen as the standard deviation of Ma1 and i, respectively. Based on these data, an in-house code of non-intrusive polynomial chaos and Sobol’ indices are employed to quantify the variability in the aerodynamic performance of a compressor cascade caused by uncertain Ma1 and i. The results show that the inflow uncertainties narrow the steady working range of the compressor cascade and lead to fluctuation in the aerodynamic performance. The fluctuation amplitude of the total pressure loss coefficient increases about 15.8 times and tenfold while the static pressure ratio increases just double and 1.3 times under various Ma1 and i conditions, respectively. An analysis of flow field under various operating conditions shows that the uncertainties change the peak isentropic Mach number on the suction surface alter the location and size of the boundary layer separation and reattachment, and shift the interference between the wake and suction surface separation vortex. The sensitivity analysis reveals that as Ma1 increases from 0.5 to 0.6, the dominant parameter affecting the uncertainty of the flow field changes from i to Ma1. When Ma1 increases to 0.75, the coupling effect of Ma1 and i dominates the flow field.

Uncertainty analysis of global and local performance impact of inflow and geometric uncertainties using sparse grid-based non-intrusive polynomial chaos

Flow variations at the inlet boundary due to the compressor operational condition changes and geometric variations of the realistic compressor blades due to the manufacturing variability cannot be absolutely avoided, the global and local performance impact of which requires to be considered in the mechanism study of performance change and the aerodynamic shape design. In this paper, a method to analyze the simultaneous impact of the inflow Mach number, inlet incidence and geometric uncertainties was proposed. To make the uncertainty modeling of geometric variations faster and closer to engineering practice, a parametric mathematical model based on scanning points on the blade was introduced to describing the profile and torsion errors in the method specially. Meanwhile, a sparse grid-based Non-Intrusive Polynomial Chaos (NIPC) was used for uncertainty quantification and uncertainty sensitivity analysis to alleviate the computational burden. Then, the method was combined with a loss source calculation method to estimate the global and local aerodynamic loss changes of a controlled diffusion compressor blade in a reference flow state of high inflow Mach number and large positive incidence, and the response performance of sparse grid-based NIPC was verified. The results show that inlet incidence and torsion error have a significance uncertainty effect on the boundary-layer separation above the suction surface, which is main reason for the fluctuation of global aerodynamic loss. The uncertainty effect of profile error on the boundary-layer separation is relatively weak, but profile error could have a certain uncertainty effect on the leading edge separation. The boundary-layer separation is insensitive to Inflow Mach number. Furthermore, a stochastic aerodynamic analysis in different reference inflow states was investigated, which reveals some laws that the uncertainty of the aerodynamic losses and the sensitivity of the inflow and geometric uncertainties change with reference inflow states.

Effects of Forebody on Supersonic Parachute Aerodynamic Performance

As a typical aerodynamic deceleration device, supersonic parachute is widely used in aerospace return and recovery. The influence of the precursor on the aerodynamic performance of parachute is very important due to operating conditions under the wake of the precursor.In order to study the effects of supersonic forebody wake on parachute performance, numerical simulation of supersonic parachute is carried out, and the calculation results are in good agreement with the results of the wind tunnel test. The flow characteristics and flow field distribution of supersonic parachutes under free flow are analyzed. On this basis, the flow field characteristics and distribution laws of forebody and parachutes under different inflow Mach numbers and the aerodynamic characteristics of parachutes are studied. The results show that the compression wave and expansion wave generated by the forebody are reflected through the wall, and the reflected wave alternately appears in the downstream flow field. Affected by the forebody wake and reflected waves, the inflow velocity of the parachute has been asymmetrically distributed, and the velocity in the upper area has been seriously attenuated.

Spiral instability modes on rotating cones in high-Reynolds number axial flow

This work shows the behavior of an unstable boundary-layer on rotating cones in high-speed flow conditions: high Reynolds number Rel&amp;gt;106, low rotational speed ratio S&amp;lt;1–1.5, and inflow Mach number M = 0.5. These conditions are most-commonly encountered on rotating aeroengine nose cones of transonic cruise aircraft. Although it has been addressed in several past studies, the boundary-layer instability on rotating cones remains to be explored in high-speed inflow regimes. This work uses infrared-thermography with a proper orthogonal decomposition approach to detect instability-induced flow structures by measuring their thermal footprints on rotating cones in high-speed inflow. The observed surface temperature patterns show that the boundary-layer instability induces spiral modes on rotating cones, which closely resemble the thermal footprints of the spiral vortices observed in past studies at low-speed flow conditions: Rel&amp;lt;105, S &amp;gt; 1, and M≈0. Three cones with half-cone angles ψ=15°, 30°, and 40° are tested. For a given cone, the Reynolds number relating to the maximum amplification of the spiral vortices is found to follow an exponential relation with the rotational speed ratio S, extending from the low- to high-speed regime. At a given rotational speed ratio S, the spiral vortex angle appears to be as expected from the low-speed studies, irrespective of the half-cone angle.

Intelligent reconstruction of the flow field in a supersonic combustor based on deep learning

The data-driven intelligent reconstruction of a flow field in a supersonic combustor aids the real-time monitoring of wave system evolution in a scramjet flow field structure, allowing the determination of the combustion state for active flow control. In this paper, a deep learning architecture based on a multi-branch fusion convolutional neural network (MBFCNN) is proposed to reconstruct the flow field in a supersonic combustor. Experiments on hydrogen-fueled scramjets with different equivalence ratios were carried out in a direct-connected supersonic pulse combustion wind tunnel with an inflow Mach number of 2.5 to establish a dataset for MBFCNN network training and testing. The trained model successfully reconstructed the flow field structure from measured wall pressure data. The flow field reconstruction model provided a rich information source for the evolution of the wave system structure under the self-ignition conditions of the hydrogen-fueled scramjet, greatly improving the detection accuracy. The proposed deep learning architecture method was compared with basic convolutional neural network and symmetric convolutional neural network methods. The three methods all accurately reconstructed the flow field of the supersonic combustor. However, the proposed MBFCNN provided the best reconstruction results, and its average linear correlation coefficient in the test set was 0.952. The proposed MBFCNN had a lower mean square error and higher peak signal-to-noise ratio than the other two methods, which verified that the proposed model is eminently able to reconstruct and predict the flow field of a supersonic combustor.