High Mach Number Research Articles

Effects of Synthetic Jets on Swirl Inflow in a Variable-Geometry Twin Air-Intake

Air intakes are an integral part of contemporary passenger and military aircraft engines. Their impact on aerodynamic performance across the entire flight envelope is critical to aircraft flight safety, efficiency, and manoeuvrability, especially at high Mach numbers due to shock waves. The high demand for reductions in aircraft weight and size and enhancements in durability, comfort, and thermal and radar signatures compel researchers and engineers to explore new designs and develop efficient air intakes for high-performance aircraft engines. Although a number of studies on air intake have been conducted and reported in the open literature, there is little information available in the public domain on bifurcated twin air intakes using synthetic jet. As a result, the primary goal of this research is to use computational fluid dynamics modelling to investigate the effects of synthetic jets on swirl inflow variable geometry twin air intake aerodynamic performance over a range of Reynolds numbers. Some important parameters (distortion coefficient, non-uniformity index, swirl coefficient, and static and total pressure coefficients) were investigated. Both static and total pressure recovery have been increased at all swirl numbers. A significant decrease in distortion coefficient and swirl coefficient has also been achieved, reaching a 53% reduction in the distortion coefficient and a 62% reduction in the swirl coefficient. The reduction in the non-uniformity index is achieved by 62% for the controlled flow case. The findings show that synthetic jets are effective in controlling the flow separation in the twin air intakes and enhancing aerodynamic performance.

High temperature non-equilibrium flow characteristics of impinging shock/flat-plate turbulent boundary layer interaction at Mach 8.42

There are fewer reports on the impinging shock/boundary layer interaction in the high Mach number and high-temperature flow than that in the supersonic flow. High-temperature flow characteristics of the impinging shock/flat-plate turbulent boundary layer interaction (IS/FTBLI) at Mach 8.42 are numerically investigated by solving two-dimensional Reynolds averaged Navier–Stokes equations coupling with the thermal–chemical non-equilibrium model. An impinging shock is formed by the wedge with a 10° deflection angle. The inviscid flow parameters ahead of the cowl of a Mach 12 inlet are selected as the free-stream condition of this study. The primary emphasis of this study lies in understanding the thermal–chemical non-equilibrium effects in the IS/FTBLI. Moreover, the chemical non-equilibrium effects similar to previous reports from others are utilized for the comparative analysis. Our findings reveal that the vibrational or thermal non-equilibrium effects exhibit maximum prominence subsequent to the intersection of the impinging shock with separation shock, as well as in the convergence area of compression waves during the flow reattachment. On the other hand, the chemical non-equilibrium effects predominantly result from oxygen dissociation and atomic nitrogen production within the boundary layer; the chemical reactions are most intense within the separation zone. By comparing with a thermally perfect gas, a reduction in the flow separation is observed in the chemical non-equilibrium effects, but the flow separation is enhanced in the thermal–chemical non-equilibrium effects. The insights gained from our research are expected to contribute to the development of flow control technology in hypersonic IS/FTBLI scenarios and aid in configuring wave structures in the inner compression section of high Mach number scramjet inlets.

Electron Heating in Shocks: Statistics and Comparison

AbstractSupernova remnant (SNR) shocks are the highest Mach number non‐relativistic shocks in electron‐ion plasmas. These shocks are the most efficient particle accelerators in space. SNR shock parameters are inferred from measurements of electromagnetic radiation from heated and accelerated particles. Temperature of the shock heated electrons is one of the most important parameters in supernova remnant shocks. Knowledge of the downstream electron‐to‐ion temperature ratio or of the ratio of the downstream electron temperature to the incident ion energy is crucial for understanding physics of the very high‐Mach number SNR shocks. Heliospheric shocks have substantially lower Mach numbers than SNR shocks but can be extensively studied in in situ observations with further extrapolation of the findings to higher Mach numbers. Magnetospheric Multiscale mission observations of the Earth bow shock are used to analyze dependence of the electron heating on the shock Mach number. It is found that the ratio of the downstream electron temperature to the incident ion energy decreases with the increase of the Mach number. At high Mach numbers this ratio and stabilizes at about 2.5%. The electron‐to‐ion temperature ratio stabilizes at about 10%. The peak electron temperature occurs at the overshoot maximum, further downstream electrons cool down. The mean ratio of the 4.5 s averages of the downstream and maximum electron temperatures is 0.85. Electron heating does not follow the thermodynamic adiabatic law. The heating and cooling behavior implies that the energy is provided by the overall cross‐shock potential while small‐scale electric fields rapidly isotropize the electron distribution.

Numerical Investigation of Double Transonic Dip Behaviors in Supercritical Airfoil Flutter

This numerical study examines the behavior of the double transonic dip with a supercritical airfoil. A two-dimensional model of a wing section with two degrees of freedom was used for the investigation. The flowfield was modeled by solving the unsteady Reynolds-averaged compressible Navier–Stokes equations using the Spalart–Allmaras turbulence model, and the equations of motion were solved to determine the structural dynamics. The present model successfully captured the behavior of the double transonic dip on the flutter boundary of the supercritical airfoil, which contrasts with the well-known behavior of the single transonic dip exhibited by conventional symmetric airfoils. Although the mechanism of the first dip at a lower Mach number corresponded to that of the well-known conventional transonic dip, the second dip at a higher Mach number was uniquely observed for the supercritical airfoil. The analysis established that, for the supercritical airfoil, the motion of the shock wave over the upper surface was significantly affected by the behavior of the boundary layer around the highly cambered aft region of the lower surface during flutter. The behavior of the boundary layer involving the separation and reattachment over the lower surface caused the unusual shock wave motion over the upper surface under the Mach number condition at the bottom of the second dip. This motion exerted negative damping forces on the motion of the airfoil, thereby becoming the primary contributor to generating the second dip experienced by the supercritical airfoil.

Dynamic multi-objective optimization of scramjet inlet based on small-sample Kriging model

The generic inlet is depicted based on a smooth Bézier curve, and the results and insights from high-dimensional dynamic multi-objective optimization of small-sample high Mach number axisymmetric scramjet inlets are discussed in detail. The optimization is performed by integrating a Kriging surrogate model-assisted improved congestion distance multi-objective particle swarm optimization algorithm and computational fluid dynamics simulation. The steady-state flow field is derived by solving the Euler equation using self-developed hypersonic internal and external flow coupling numerical simulation software, which is designed to minimize inlet surface area and drag while improving the total pressure recovery factor. The results revealed that the generic inlet can achieve a total pressure recovery capability exceeding 95%, with minimal surface area and drag. The prediction error, mean absolute percentage error, of the performance dynamic surrogate model based on Kriging is less than 1%, and the performance parameter optimization shows an improvement greater than 8% compared to static multi-objective optimization results. Ultimately, the obtained Pareto solution set is grouped by K-means feature recognition, contributing to a comprehensive understanding of the flow physics knowledge related to optimal geometric local shape control. Finally, an inward-turning inlet is designed by streamline tracking technology based on the optimized axisymmetric scramjet inlet primary flow field.

STAJe−: A Supersonic Flight Program STAJe− 1 Payload Design and System Study

The STAJe− flight program is the latest implementation of a university-based supersonic and hypersonic flight-test program. In the first flight, which is the subject of this study, the focus is on evaluating the suitability of commercial off-the-shelf electronic hardware, additively manufactured internal payload structures, and a long-range communications technique in the challenging environment presented by flight at high supersonic Mach numbers. These items can significantly reduce the cost of a flight experiment, and thus lower the threshold for flight experimentation. The scientific objectives of this flight are to collect pressure and temperature data on the whole trajectory to compare to numerical models, to optically measure the temperature of composite fins mounted on the payload, and to eject the flight module from the rocket assembly and track its location for subsequent recovery. Design calculations for the tracking system are presented alongside experimental results from ground tests demonstrating the suitability of the system. Furthermore, a one-dimensional calculation of the external wall temperatures over the trajectory is presented, which shows that internal systems are sufficiently protected from the harsh environment. Although the materials are shown to be appropriate for the flight, a layer of cork will be added to parts of the payload to further reduce the chance of structural failure. In addition, it is shown that a pneumatic ejection system is suited to overcoming the drag forces on the front of the rocket to initiate separation.

Parametric optimization and exergy analysis of a high mach number aeroengine with an ammonia mass injection pre-compressor cooling cycle

The Ammonia mass injection pre-compressor cooling (Ammonia MIPCC) cycle provides a substantial potential for the development of a high Mach aeroengine. To evaluate performance and analyze exergy of Ammonia MIPCC aeroengine, the parameter matching modeling method and Sequential Quadratic Programming (SQP) optimization are applied. These proposed methods can realize the constraints of compressor outlet temperature, combustor outlet temperature, power balance and flow conservation, thereby optimizing the specific impulse performance. Results show that the optimal operating line (OOL) of the Ammonia MIPCC aeroengine can be divided into two paths: the temperature limit operating line and the maximum specific impulse operating line. Moreover, the OOL of the Ammonia MIPCC aeroengine is always in a high exergy utilization state. When the ammonia rate increases from 0 to 0.08, the changes in exergy reduction rate, specific impulse and specific thrust corresponding to OOL at Ma 3.0 are 18.89%–31.83%, 2347.47 to 1113.56s, 768.91–953.47 N/(kg/s), respectively. Additionally, results show that the cycle with lower inflow cooling requirements can improve economy. In summary, parameter optimization and exergy analysis help to determine the working characteristics of high Mach number Ammonia MIPCC aeroengines, enhance the engine thrust performance and extend the operating envelope.

Particle Energization at a High Mach Number Perpendicular Shock: 1D Particle-in-cell Simulations

In this paper, we use a 1D particle-in-cell simulation code to study particle preaccelerations at a high Mach number perpendicular shock. Our simulation results show that almost all of the injected particles can be reflected at the shock front, and then they immediately gyrate back to upstream for a long distance. That facilitates the formation of a large-scale shock foot where they dominate the average velocity of particles and the formation of resultant electric fields with several subareas, unlike a low Mach number shock with fewer reflected particles. In the large-scale shock foot with subareas, these reflected particles can be energized by the motional electric fields and unexpected electrostatic fields, which means they may undergo multiple stages of preacceleration processes when gyrating just before the high Mach number perpendicular shock front with high-intensity particle reflection.

Enhancing the mixing characteristics of multi-hydrogen jets in scramjet engines through the implementation of fuel injection strategies and vortex generator positioning

Fuel injection strategies and vortex generators (VGs) play vital roles in effectively managing temperature distribution within the combustion chamber of scramjet engines. By optimizing the fuel injection mechanism and strategically positioning VGs, the efficiency of fuel/air mixing can be significantly improved. In this study, k-epsilon turbulence model used to accurately capture shock waves. The operating Mach number was maintained at 2.0 through the use of a pressure far-field inlet. The investigation focused on three different fuel injection mechanisms: backward-facing, transverse, and forward-facing, while also examining the mixing capabilities of VGs located at varying distances from the inlet. Detailed analysis was conducted to understand the impact of these factors. Backward flow injection was found to promote efficient mixing and reduce the risk of flame blowout. Furthermore, optimizing the injection location proved instrumental in enhancing fuel and air mixing efficiency, leading to higher Mach numbers and improved combustion efficiency. However, it is important to avoid excessively high Mach numbers as they can induce flame instability and blowout. Among the different injection methods, multiple port injection demonstrated the highest Mach numbers, while backward-facing injection yielded the lowest. Multiple port injection also achieved high static pressure. Introducing VGs into the system significantly increased static pressure, and optimal fuel injection and VGs location resulted in stable combustion, improved combustion efficiency, and minimized unburned fuel. The injection method and the presence of VGs greatly influenced turbulence kinetic energy. Notably, VGs located at distances of 81 mm and 31 mm from the inlet produced the highest turbulence intensity for forward-facing injection and transverse fuel injection, respectively.

On the Over-Prediction of Centrifugal Compressor Pressure Ratio in the High Impeller Tip Mach Number Regime

Abstract Reynolds-averaged Navier–Stokes (RANS) simulation is a routinely used tool for turbomachinery research and development, but it often over-predicts the pressure ratio of centrifugal compressors especially in the high impeller tip Mach number regime. In this work, the effects of a series of geometrical and numerical uncertainties and errors on the aerodynamic performance of a centrifugal compressor are investigated systematically. The investigated compressor is the NASA CC3 centrifugal compressor under different impeller tip Mach number conditions. The investigated geometrical/numerical factors include the impeller blade fillet, the impeller hub cavity, the impeller running tip clearance, the averaging method in the post-process, the turbulence model, the inlet duct hub rotation, and the inlet turbulence boundary condition. Results show that the uncertainty of the predicted total pressure ratio generally increases with the impeller tip Mach number. Among the investigated factors, the averaging method, the impeller blade fillet, the turbulence model, and the impeller hub cavity have the most pronounced effects in determining the compressor total pressure ratio. By adopting realistic geometric features, advanced turbulence modeling treatments, and the same averaging method as the experiment, the over-prediction in the total pressure ratio can be alleviated. Detailed flow mechanism analysis with respect to the impeller hub cavity and the turbulence model has been performed. These findings provide valuable guidance for future RANS simulations of centrifugal compressors.

Performance analysis of a thermal management system based on hydrocarbon-fuel regenerative cooling technology for scramjets

Thermal protection and power generation are two of the major technical challenges in developing hypersonic vehicles with long-range/endurance. This research is aimed to provide a system for thermal protection and power generation for hydrocarbon-fueled scramjets. This study proposed a thermal management system based on fuel vapor turbine (FVT) and closed-Brayton-cycle (CBC) to generate power, and a bypass loop is connected with CBC in parallel to ensure thermal protection demand. A performance analysis model considering the combustor and cooling channels is established. The combustion and heat transfer process are described in a quasi-one-dimensional way, while FVT and CBC are described in a zero-dimensional form. Results indicate that large CBC power generation efficiency (ηPG,CBC) and FVT expansion ratio (πFVT) can improve power generation with the same thermal protection performance. Because of the integrated variation law of cooling heat exchange and ηPG,CBC, there exists an optimal value of fuel outlet temperature (Tf3) while meeting the thermal protection requirement. The high Mach number and low equivalence ratio both require a significant increase in the cooling capability of fuel, and it is feasible to sacrifice the power generation of the CBC subsystem to meet the thermal protection demand.

Effect of Flare Angle and Natural Ventilation on The Aerodynamic Characteristics of a Typical Re-Entry Body At Subsonic and Transonic Mach Numbers

Aerodynamic force and moment measurements were carried out on models of a typical re-entry body, featuring a blunt nose, conical body and flare, to assess the effect of flare angle on the overall aerodynamic characteristics. In one of the configurations, the junction between the cone and the flare was vented out to the base of the model to assess the effect of ventilation on the aerodynamic characteristics. It was expected that the stability characteristics would improve, by virtue of improved contribution from the flare to the normal force of the configuration. However, measurements did not show the expected trends. Instead, there was a remarkable reduction in the axial force with ventilation, by as much as 25% all across the Mach number and angle of attack range. Base pressure measurements indicated that in the low Mach number range pressure rise occured. However, in the high transonic Mach number range, even though the force measurements showed drag reduction, there was no significant rise in base pressure. Ventilation showed delay in transonic drag rise. It is therefore suggested that the overall pressure field is modified by ventilation, which results in drag reduction of the configuration.

Autoignition of two-phase n-heptane/air mixtures behind an oblique shock: Insights into spray oblique detonation initiation

Autoignition of n-heptane droplet/vapor/air mixtures behind an oblique shock wave are studied, through Eulerian-Lagrangian method and a skeletal chemical mechanism. The effects of gas/liquid equivalence ratio (ER), droplet diameter, flight altitude, and Mach number on the ignition transient and chemical timescales are investigated. The results show that the ratio of chemical excitation time to ignition delay time can be used to predict the oblique detonation wave (ODW) transition mode. When the ratio is relatively high, the combustion heat release is slow and smooth transition is more likely to occur. In heterogeneous ignition, there are direct interactions between the evaporating droplets and the induction/ignition process, and the chemical explosive propensity changes accordingly. The energy absorption of evaporating droplets significantly retards the ignition of n-heptane vapor. In the two-phase n-heptane mixture autoignition process, the ignition delay time decreases exponentially with flight Mach number and increases first and then decreases with the flight altitude. As the liquid ER increases, both ignition delay time and droplet evaporation time increase. With increased droplet diameter, the ignition delay time decreases, and the evaporation time increases. Besides, when Mach number is less than 10, the ratio of the chemical excitation time to ignition delay time generally increases with the flight altitude or Mach number. It increases when the liquid ER decreases or droplet diameter increases. When the Mach number is sufficiently high, it shows limited change with fuel and inflow conditions. The ODW is more likely to be initiated with a smooth transition at high altitude or Mach number. Abrupt transition mode tends to happen when fine fuel droplets are loaded. The results from this work can provide insights into spray ODW initiation.

Thermochemical non-equilibrium flow characteristics of high Mach number inlet in a wide operation range

The high-temperature non-equilibrium effect is a novel and significant issue in the flows over a high Mach number (above Mach 8) air-breathing vehicle. Thus, this study attempts to investigate the high-temperature non-equilibrium flows of a curved compression two-dimensional scramjet inlet at Mach 8 to 12 utilizing the two-dimensional non-equilibrium RANS calculations. Notably, the thermochemical non-equilibrium gas model can predict the actual high-temperature flows, and the numerical results of the other four thermochemical gas models are only used for comparative analysis. Firstly, the thermochemical non-equilibrium flow fields and work performance of the inlet at Mach 8 to 12 are analyzed. Then, the influences of high-temperature non-equilibrium effects on the starting characteristics of the inlet are investigated. The results reveal that a large separation bubble caused by the cowl shock/lower wall boundary layer interaction appears upstream of the shoulder, at Mach 8. The separation zone size is smaller, and its location is closer to the downstream area while the thermal process changes from frozen to non-equilibrium and then to equilibrium. With the increase of inflow Mach number, the thermochemical non-equilibrium effects in the whole inlet flow field gradually strengthen, so their influences on the overall work performance of the high Mach number inlet are more obvious. The vibrational relaxation or thermal non-equilibrium effects can yield more visible influences on the inlet performance than the chemical non-equilibrium reactions. The inlet in the thermochemical non-equilibrium flow can restart more easily than that in the thermochemical frozen flow. This work should provide a basis for the design and starting ability prediction of the high Mach number inlet in the wide operation range.

A deep reinforcement learning control approach for high-performance aircraft

This research introduces a flight controller for a high-performance aircraft, able to follow randomly generated sequences of waypoints, at varying altitudes, in various types of scenarios. The study assumes a publicly available six-degree-of-freedom (6-DoF) rigid aeroplane flight dynamics model of a military fighter jet. Consolidated results in artificial intelligence and deep reinforcement learning (DRL) research are used to demonstrate the capability to make certain manoeuvres AI-based fully automatic for a high-fidelity nonlinear model of a fixed-wing aircraft. This work investigates the use of a deep deterministic policy gradient (DDPG) controller agent, based on the successful applications of the same approach to other domains. In the particular application to flight control presented here, the effort has been focused on the design of a suitable reward function used to train the agent to achieve some given navigation tasks. The trained controller is successful on highly coupled manoeuvres, including rapid sequences of turns, at both low and high flight Mach numbers, in simulations reproducing a prey–chaser dogfight scenario. Robustness to sensor noise, atmospheric disturbances, different initial flight conditions and varying reference signal shapes is also demonstrated.

Numerical investigation on the initiation of oblique detonation waves in stoichiometric methane–air mixtures

The study investigates the oblique detonation characteristics in a stoichiometric methane–air mixture employing reactive Euler equations coupled with a detailed chemical reaction model. The analysis evaluates the initiation variation for both the inflow Mach number and the wedge angle. The results show that an increase in the Mach number from M0 = 9 to M0 = 12 changes the oblique shock-to-detonation transition from an abrupt to a smooth type. Furthermore, the distinction between high and low Mach numbers is explored through a comparison of the induction length along the wave surface. Additionally, the transition is affected by changes in wedge angle elevation under high Mach numbers, while a large wedge angle generates a second detonation wave surface due to a strong reflection shock downstream under low Mach numbers. Finally, the initiation length is quantitatively compared across varying inflow Mach numbers and wedge angles, which can benefit future investigations of a methane-fueled oblique detonation engine.

Two‐Dimensional Hybrid Simulations of High‐Speed Jets Downstream of Quasi‐Parallel Shocks

AbstractHigh‐speed jets (HSJs) are frequently observed downstream of a quasi‐parallel bow shock, and they are considered to play important roles in the coupling of the solar wind, the magnetosheath, and the magnetosphere. Using two‐dimensional hybrid simulations, we study the formation of HSJs in quasi‐parallel shocks with different shock angles (θBn). The interaction of the upstream compressive structures that are inhomogeneous in the direction perpendicular to the background magnetic field and the shock front leads to the shock ripples, and then the downstream HSJs. In a parallel shock with the shock angle θBn = 0°, the interaction regions of the upstream compressive structures with the shock front don’t change with time. The shock ripples remain in the same regions of the shock front, and then the HSJs can develop into large‐scale sizes, especially under low plasma β or high Mach number (MA). While in a quasi‐parallel shock with a non‐zero shock angle, the interaction regions of the upstream compressive structures with the shock front move along the shock front. The shock ripples change with time, and the scale size of the downstream HSJs becomes smaller with the increase of the shock angle.

Energetic ion enhancements in sheaths driven by interplanetary coronal mass ejections

We analyze here an energetic proton enhancement in a sheath ahead of a slow interplanetry coronal mass ejection (ICME) detected by Parker Solar Probe on June 30, 2021 at the heliospheric distance of 0.76 AU. The shock was likely quasi-parallel and had a high Mach number. However, the proton fluxes were not enhanced at the shock but about an hour later. The fluxes stayed elevated with a sporadic behaviour throughout the sheath. We suggest that some mechanism internal to the sheath was responsible for the energization. The observations show enhanced levels of magnetic field fluctuations in the sheath and frequent presence of highly reduced magnetic helicity structures (sigma _{m}) at various time scales, representing either small-scale flux ropes or Alfvénic fluctuations that could have contributed to the energization. The correlation between the energetic proton fluxes and normalized fluctuation amplitudes/occurrence of high sigma _{m} structures was generally weak or negligible. The most striking feature of the sheath was a strong enhancement of density (up to 50 cm−3) that implies the importance of compressive acceleration in the sheath. A statistical analysis of ion enhancements of 73 sheaths detected by ACE at {sim} 1 AU reveals that this sheath was peculiar as in ICME-driven sheaths preceded by strong shocks the ion fluxes typically peak at the shock and strongly decline through the sheath.

Controlling second-mode oblique breakdown in high-speed boundary layers using streak: A direct numerical simulation study

The suppression effect of streaks on the second mode has been studied by the stability analysis in the past decade. In the present study, we conduct direct numerical simulations (DNS) to investigate the direct nonlinear control in the complete laminar-to-turbulence transition scenario within a high Mach number supersonic boundary layer. Our study aims to provide a comprehensive understanding of the mechanisms underlying the influence of streaks on the transition of high-speed boundary layers. Our work demonstrates the effective control effect of the streaks generated by blowing and suction strips on laminar-to-turbulence transition via the second-mode oblique breakdown at a Mach number 4.5 high-speed boundary layer using DNS. Modal analysis, nonlinear disturbance formulation, and stability analysis are used to provide insight into the stabilization effect of control streaks. Crucially, the role of three-dimensional control modes and mean-flow distortion generated by the control streak are investigated. Our findings indicate that both mean-flow distortion and three-dimensional control modes effectively stabilize the fundamental oblique second mode, particularly when the amplitude of control streaks is strong.

Effect of Subsonic Compressibility and Constraints on Generation of 3-D Optimal Camber Shapes

A flat surface wing is optimized for minimum induced drag for low and high subsonic Mach numbers. Wing for transport airplane role is considered. Optimal warp is separated into twist and camber. Effect of compressibility is analysed on the camber shape generated for two different Mach numbers. Constraint of lift is firstly used for developing an objective function for drag minimization. Next the lift and wing root bending moment constraints are considered to form objective function. Analysis is made on the resulting camber and twist distributions.