Pressure profiles and flame behaviors of a scramjet model combustor with two-stage fuel injection for low flight Mach numbers

In this study, a new actuator for generating a supersonic jet pulsating at an ultrasonic frequency (several tens of kHz), which is called ultrasonic-driven pulsating supersonic jet (USDPSJ) actuator, is proposed. This actuator consists of a flat wall with a 1.0-mm-diameter orifice and a top surface of an ultrasonic oscillator with a sub-millimeter gap between them. A supersonic jet pulsating at an ultrasonic frequency is successfully generated by driving the oscillator under the conditions that the ratio of the pressure upstream to downstream of the orifice is higher than the ratio at which choking occurs in the flow passing through the orifice. The jet is visualized by a schlieren method using a high-speed video camera. The visualization results reveal that a pulsating supersonic jet with a shock cell structure is generated. Furthermore, it is found that the jet is generated intermittently under specific conditions regarding the gap height, oscillator amplitude, and pressure ratio. The time-varying jet Mach numbers are calculated from the periodically fluctuating pressures measured by a pressure sensor placed at the orifice location under the same oscillator driving conditions as when the jets are generated from the orifice. The calculated jet Mach numbers are confirmed to agree reasonably well with the jet Mach numbers estimated from the shock cell lengths of the visualized jets. Based on this confirmation, the time-varying jet velocities are estimated. The estimation results reveal that the jet pulsates over its velocity fluctuation range from 0 to above 500 m/s under specific conditions regarding the gap height and oscillator amplitude, when the pressure ratio exceeds 2, demonstrating that the proposed actuator can achieve unprecedentedly large velocity fluctuations at an ultrasonic frequency. In addition, based on the experimental results, the maximum values for the actuator's energy conversion efficiency, jet penetration distance, and jet spreading rate are estimated to be 45%, 15.1 times the orifice diameter, and 0.35, respectively. • A new actuator is proposed that can create a pulsating supersonic jet at an ultrasonic frequency. • The proposed actuator is driven by an ultrasonic transducer. • The proposed actuator can generate a pulsating jet over its velocity fluctuation range of 0 to above 500 m/s. • The proposed actuator requires the pressure only 2 times higher than the back pressure to achieve the range of 0 to above 500 m/s.

Low Mach number compressible multiphase particle-in-cell method for viscous flow problem

Accurate and rapid aerodynamic prediction is essential for projectile trajectory simulation and digital twin driven exterior ballistic systems, where conventional CFD methods are computationally expensive and difficult to deploy in real-time applications. To address this challenge, this paper proposes a primary-residual dual-stage surrogate modeling method based on a multilayer perceptron (MLP). A primary prediction model for aerodynamic parameters is constructed using CFD simulation data, where Mach number and angle of attack are used as input variables to establish the global nonlinear aerodynamic mapping. By introducing a residual learning mechanism, a compensation model is established to address fitting deviations in highly nonlinear regions, thereby correcting the nonlinear biases of the primary model predictions. The residual network further captures localized nonlinear discrepancies not fully resolved by the primary model, significantly improving prediction accuracy in high-Mach-number and large-angle-of-attack regions. Comparative validation against polynomial fitting, RBF, and single-stage MLP models shows that the proposed method achieves the best overall performance. For example, the RMSE is reduced from 2.91 to 0.87 for lift prediction and from 0.634 to 0.086 for pitching moment prediction. Overall, the proposed model reduces the average RMSE by approximately 50% across all aerodynamic parameters, demonstrating stronger robustness, improved generalization capability, and practical engineering applicability for digital twin based projectile exterior ballistic prediction.

Abstract We analyze the high velocity cloud HVC 53.9+1.3 using the highly sensitive and high angular resolution HI data piggybackly recorded by the FAST Galactic Plane Pulsar Snapshot (GPPS) survey. Detailed sub-structures of the HVC are resolved, the densest peak of the HVC has a HI column density of N HI = 3.5 × 10 19 cm -2 , two orders lower than the main component of the Galactic HI disk. Using the Gaussian decomposition method, cold and warm phases of HI gas clouds are separated. Cold HI gas is found in a supersonic state with Mach number of M = 2.3, and tends to associate with high HI density, with cold cores surrounded by skins of warmer gas, while warmer gas are in subsonic state with M < 1.

The dynamics of detonation is governed by intrinsic longitudinal and transverse instabilities. This work fills the research gap of detonation stability at elevated pressure as real gas effects, including the finite molecular volume and inter-molecular attraction forces, become non-negligible. The stability of one-dimensional pulsating and two-dimensional cellular detonations at elevated pressure was investigated using both linear stability analysis and high-fidelity numerical simulations. For pulsating detonations, the neutral stability boundaries derived from the linear stability analysis reveal that the impact of real gas effects on detonation stability depends on the heat content. Specifically, real gas effects stabilise pulsating detonation at high heat content, destabilise it at intermediate levels and exhibit competing influences, i.e. destabilisation from finite molecular volume and stabilisation from the inter-molecular attraction force, at low heat content. For cellular detonations, a stability map is proposed in the Mach number–reduced activation energy plane, demonstrating that real gas effects stabilise detonation and lead to more regular cellular structures or even fully planar detonations. Theoretical predictions show good agreement with numerical simulations. An oscillator model was used to elucidate the interplay between the heat release and shock dynamics. It shows that the stabilising effect of real gas is associated with reduced velocity oscillation amplitude, increased phase lag between heat release and shock speed variations and lower values of the coupling parameter. The comprehensive linear stability analysis with complete real gas models is an effective predictive tool for the detonation dynamics at elevated pressure.

We present a perturbation-based framework that captures buoyancy effects on modal instabilities in stratified boundary-layer flows within the fully compressible, non-Oberbeck–Boussinesq formulation. Treating the Richardson number as a small parameter and recasting the stability problem into an adjoint-residual form, we derive a first-order correction for the eigenvalues using only the neutrally buoyant eigenvalue problem. The framework applies to both ideal-gas and non-ideal fluid boundary layers and eliminates the need to re-solve the eigenvalue problem at each stratification level at minimal computational cost. For ideal-gas boundary layers, the framework accurately predicts how stable and unstable stratification modifies Tollmien–Schlichting waves, from growth rates and eigenfunctions to $N$ -factors, across a wide range of Prandtl numbers, temperature ratios and Mach numbers. Notably, the buoyancy sensitivity varies strongly with Prandtl number, revealing that for a given Richardson number, buoyancy can switch from destabilising to stabilising depending on the fluid. Beyond ideal-gas conditions, we apply the first-order buoyancy correction to strongly stratified boundary layers with supercritical fluids, where the phase relationship between density and velocity perturbations determines whether buoyancy stabilises or destabilises the underlying instability. The resulting $N$ -factors demonstrate, for the first time, that buoyancy significantly affects transition predictions under pseudo-boiling conditions.

We investigate the influence of realistic sheared grazing flow on acoustic propagation in three-dimensional rectangular ducts. We show that the conclusions reached in the literature about the effects of sheared grazing flow on acoustic propagation in lined ducts are dependent on the flow profiles used in those studies and that significantly different conclusions are reached once a realistic flow profile is used. We particularly focus on small ducts typical of most experimental impedance eduction facilities, for which velocity gradients are relevant in a significant fraction of the duct cross section. We assess the effect of simplifying the velocity distribution in the cross section to either a one-dimensional (two-dimensional, spanwise-infinite duct) or a uniform flow profile. Three flow profiles are considered, namely, i) the tensorized hyperbolic tangent, ii) the law of the wall, and iii) one obtained from a RANS simulation. These flow profiles are used as input in numerical simulations, based on the solution of the three-dimensional Pridmore–Brown equation, to perform in silico impedance eduction experiments. Results show that realistic flow profiles can be well approximated for acoustic wave propagation in ducts by uniform or one-dimensional flow profiles, provided that the bulk Mach number is correctly accounted for, which contrasts with previous findings considering more simplistic flow profiles. The key conclusion of this work is that if viscous effects are negligible and acoustic impedance is a good representation of a lined wall with grazing flow, then the simplification to a uniform flow is a reasonable approximation, and traditional eduction methods are sufficiently accurate.

Conical shock wave interacting with the boundary layer on a flat plate (Conical SBLI) is a complex three-dimensional phenomenon distinguished by a hyperbolic shock trace and spanwise variation in interaction intensity. This study investigates upstream-influence behavior in a conical shock wave and turbulent boundary layer interaction using numerical simulations at a Mach number of 2 and a unit Reynolds number of 30×106/m. The simulations were performed for cone-half angles ranging from 14–30 deg, and results were validated against existing experimental data. The analysis reveals that the upstream influence length on the symmetry plane scales with the pressure rise, following a power law similar to that observed in two-dimensional (2D) shock wave and boundary layer interactions. Spanwise measurements show that the normal upstream influence length increases along the span, transitioning from cylindrical similarity near the symmetry plane to a conical similarity further outboard. These findings provide new insights into conical shock wave and turbulent boundary layer interaction, and offer a framework for predicting upstream influence length in practical supersonic flows over conical geometries.

This study experimentally investigates the combustion mode transition and oscillation phenomena during hydrogen-ethylene fuel switching in a cavity-stabilized combustor with the inflow Mach number of 2.52. Reduced-order variational mode decomposition (RVMD) was used to identify the flame characteristics during mode transition in the upstream hydrogen-ethylene switching condition. The results reveal that no distinct mode transition is observed for the downstream hydrogen jet. In contrast, the upstream hydrogen jet significantly enhances combustion intensity and combustor pressure during the switching process, leading to a transition to the wake combustion mode. It is found that flame oscillations weaken during the hydrogen-ethylene co-combustion stage, whereas large-amplitude oscillations emerge after the switching process. This is attributed to the precombustion of upstream hydrogen during the switch to ethylene, which generates high-temperature zones and recirculation regions between these two jets. The enhanced combustion provides sufficient preheating to promote ethylene flame propagation, enabling the transition to the wake combustion mode. After hydrogen has been turned off, the disappearance of upstream recirculation weakens the ethylene flame, causing mismatched flame-flow interactions and resulting in significant oscillations.

The physical mechanism of oscillatory instability in spiked-body flow at a freestream Mach number of 6 and a Reynolds number of 0.13 million, based on the cylinder diameter, is investigated using high-fidelity delayed detached-eddy simulation (DDES). The simulation successfully captures the unsteady shock structures and global flow features, enabling a detailed analysis of the underlying dynamics. A dominant low-frequency mode governing the large-scale oscillation is identified and shown to be intrinsically linked to the evolution of the complex coupled interaction among shock waves, shear layers, and separation regions. The oscillation cycle is characterized by the periodic advance and retreat of the separation zone, accompanied by the flapping motion of the separation shock and the shear layer. The bidirectional propagation of inner compression waves and their convection velocities are identified, revealing their key role in sustaining the feedback loop among the separation shock, shear layer, and recirculation region. These findings provide a new physical understanding of the origin and sustaining mechanism of large-scale oscillations in hypersonic spiked-body flows.

Thermodynamic coupling mechanisms in precooled turbine-based combined cycle: From fuel properties to multi-objective performance characterization

A scaling framework unifying the markedly different and independently studied cylindrical and spherical shock convergence is presented. For ionizing argon, we show that the focal temperature becomes invariant to shock symmetry and initial shock conditions when scaled by the prefocus shock Mach number, after accounting for pressure effects. The resulting collapsed focal temperature-Mach number relation is governed by the thermodynamics of argon, up to an equilibrium temperature of 35 000K. Such a scaling enables both predictive estimation of focal temperatures over a wide range of initial conditions and, conversely, determination of the parameters required to achieve a target temperature in a given medium.

A molecular-kinetic hypothesis on the mechanics of compressible gas flow at low Mach numbers

ABSTRACT The dynamic stabilization of detonation waves is critically governed by two factors: the incoming flow Mach number and the combustor geometry. This study investigates their combined role using high-fidelity numerical simulation. We specifically analyze a configuration featuring a cavity and an expansion wall. The objective is to elucidate the mechanisms through which these geometric elements interact with the flow to promote sustained detonation above the Chapman-Jouguet (CJ) velocity. The two-dimensional reactive Navier – Stokes equations coupled with a one-step, two-species reaction model are solved using a hybrid sixth-order Weighted Essentially Non-Oscillatory Central Difference (WENO-CD) scheme within a Structured Adaptive Mesh Refinement (SAMR) framework. This high-fidelity computational approach captures the essential shock waves, contact discontinuities, and fine-scale reaction zones with high resolution and minimal numerical dissipation, making it well-suited for simulating the intricate physics of detonation stabilization. The results demonstrate that a combustor configuration incorporating a cavity and an expansion wall promotes dynamic detonation stability. Specifically, the Prandtl – Meyer expansion fan and large-scale vortices generated by the expanding wall, along with pressure oscillations from the cavity, enhance the consumption and heat release of the unburned jet, thereby sustaining detonation. The incoming Mach number is found to critically influence stability: detonation waves cannot stabilize below the Chapman – Jouguet (CJ) velocity, but can be stabilized above it through the combined effects of the cavity and expansion wall. Additionally, geometric parameters such as the expansion wall angle and cavity dimensions significantly affect the stability and location of the detonation wave. Increasing the expansion angle extends the unburned jet and attenuates detonation, while altering cavity size and position influences pressure oscillations and combustion efficiency, thereby modulating the duration and degree of wave stability. This highlights the need for careful geometric optimization to achieve desired operational characteristics, balancing the reinforcement mechanisms from the cavity with the flow-turning effects of the expansion.

Density fluctuations in the shear layer locally alter the effective index of refraction of the atmosphere, causing bore-sight errors that are characterized by an apparent shift in the target location. To address the lack of viable correction methods for supersonic and hypersonic aero-optical distortions, we perform a large-eddy simulation using the JENRE Multiphysics Framework to approximate the boundary-layer and shear-layer flow over a cavity operating at a free-stream Mach number of 2.3 and an altitude of 16km. The optical path difference (OPD) is calculated from the high-frequency density sampling over a 0.0254m×0.0254m aperture located at the center of the cavity. Spectral proper orthogonal decomposition of the OPD reveals dominant flow structures contributing to wavefront aberrations. Using the simulated OPD data, we train an artificial neural network to process the Shack-Hartmann wavefront sensor outputs and reconstruct the original wavefront. This data-driven approach demonstrates potential for faster and more accurate correction of imaging errors compared to traditional methods, particularly when tailored to specific operational conditions.

Modeling atmospheric and stellar phenomena requires understanding compressive turbulence, a more complex problem than its incompressible counterpart. This paper employs a novel mathematical framework to analyze energy transfers and fluxes in subsonic compressible flows. We perform direct numerical simulations on a 10243 grid for turbulent Mach numbers 0.15, 0.30, and 0.45. We apply stochastic random forcing to both rotational and compressive modes. We demonstrate that for subsonic flows, energy transfers from solenoidal to compressive modes are confined primarily to large scales, allowing independent rotational and compressive kinetic energy cascades. Consequently, both components maintain constant inertial-range fluxes, resulting in Kolmogorov scaling for the rotational velocity and Burgers scaling for the compressive velocity. We also observe that compressive kinetic energy is converted to internal energy via pressure dilatation. This advancement enables further exploration of locality, compressible convection, and compressible magnetohydrodynamics.

Abstract Single-pointing observations of 73 Planck cores from the Early Cold Core Catalogue with the Nanshan 26-m telescope are presented targeting the H2CO (11, 0–11, 1) 4.8 GHz (λ ~ 6 cm) absorption line. H2CO absorption has been detected in 51 sources (69.9%), with 24 (32.9%) also showing components with hyperfine structure (HFS). In these 51 detected cores, non-thermal velocity dispersion dominates over thermal line broadening (σTH/σNT < 1), with 96% exhibiting supersonic turbulence ($\mathcal {M} > 1$). A weak correlation between σNT and Tkin suggests that turbulence contributes to gas heating. A strong σNT – ortho-H2CO column density correlation highlights the importance of both turbulence and gravity. For the 24 sources with resolved HFS, the derived excitation temperatures ranges from 2.08 to 2.59 K (mean 2.37 K). Follow-up mapping of four high-S/N cores with regions of resolved HFS components reveals widespread gas with Tex ≈ 2.36–2.64 K. Cores with resolved HFS exhibit narrower line widths, lower Mach numbers, higher column densities, and larger optical depths, indicating dynamically quiescent gas in clouds moving towards early gravitational collapse. In contrast, non-HFS regions surrounding the HFS regions in these sources display broader lines and stronger non-thermal motions, suggesting a dynamically complex environment where gravity begins to influence the earliest stages of star formation.

Turbojet engines remain critical propulsion systems for high-speed aircraft, where performance optimization is essential to improve thrust efficiency and operational capability under varying flight conditions. This study presents a performance optimization framework for a turbojet engine using Particle Swarm Optimization (PSO). A thermodynamic mathematical model of the turbojet cycle was developed to evaluate engine performance as a function of compressor pressure ratio, turbine inlet temperature, Mach number, and flight altitude. The PSO algorithm was employed to determine the optimal operating parameters that maximize engine thrust while considering realistic operational constraints. Simulation results indicate that the optimized configuration achieved a compressor pressure ratio of 9.27, turbine inlet temperature of 1800 K, Mach number of 0.90, and altitude of 0 m. Under these conditions, the engine produced a maximum thrust of approximately 152 KN, representing a performance improvement of about 235% compared to the baseline thrust of 45.4 kN. The optimization results were further validated using a MATLAB Simulink implementation of the turbojet engine model, which produced a thrust of approximately 160 kN, confirming the reliability of the optimization framework. The results demonstrate that PSO provides an effective and robust approach for turbojet engine performance optimization under variable flight conditions.

The near-ground flight of new-generation aerospace vehicles—such as horizontally launched electromagnetic systems and sea-skimming missiles—requires a comprehensive understanding of high-speed ground effect (GE) aerodynamics. However, transonic and supersonic GE are far less studied than the subsonic regime. This work systematically examines GE aerodynamics of a representative airfoil across subsonic, transonic and supersonic regimes and a wide range of flight heights by numerically solving the compressible Reynolds-Averaged Navier–Stokes equations. The evolution of flow patterns and aerodynamic characteristics with Mach number in GE is provided to complement classical textbooks focusing on unbounded flows. Relative to subsonic GE, aerodynamic force variations in the transonic and supersonic regimes are substantially larger. The conventional critical height criterion of h/c≈1 becomes invalid at higher Mach numbers, where significant GE persists even at elevated heights. As the height decreases, the lower critical Mach number increases slightly, whereas the upper critical Mach number rises sharply. Across all regimes, the total lift is primarily governed by pressure changes on the lower surface, although the dominant mechanisms differ: in the subsonic regime, the blocking and camber effects prevail, whereas in the transonic and supersonic regimes, the complex shock-ground interactions are the principal contributors.