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On 18 November 2023, SpaceX launched the Starship, the tallest and the most powerful rocket ever built. The Super Heavy engine separated from the Starship spacecraft and exploded at 90 km of altitude, while the main core Starship continued to rise up to 149 km and exploded after ∼8 min of flight. In this work, we used data from ground‐based GNSS receivers and we analyzed total electron content (TEC) response to the Starship flight and the two explosions. For the first time, we observed large‐distance northward propagation of intensive 2,000 km V‐shaped ionospheric disturbances from the rocket trajectory. The observed perturbations, most likely, represent shock waves propagating with the cone angle of ∼14° on the North and ∼7° on the South against the flight track that corresponds to the Mach angle of the shock waves in the lower atmosphere. The Starship explosion also produced a non‐chemical depletion in the ionospheric TEC. Plain Language Summary: On 18 November 2023, SpaceX launched the Starship, the tallest and the most powerful rocket ever built. About 2 min and 40 s after the liftoff, the Super Heavy engine separated from the Starship spacecraft and exploded at an altitude of 90 km. The main core Starship continued to rise to 149 km and exploded as well. The rocket launch and explosion produced an unexpected response in the ionosphere—the ionized part of the Earth's atmosphere. The Starship flew at a velocity, exceeding the local sound speed, and generated cone‐like atmospheric shock‐acoustic waves. Most unexpectedly, the observed disturbances represented long and intensive multi‐oscillation wave structures that propagated northward, which is unusual for disturbances driven by a rocket launch. The Starship explosion also generated a large‐amplitude total electron content depletion that could have been reinforced by the impact of the spacecraft's fuel exhaust in the lower atmosphere. This study appears to be the first‐time detection of a non‐chemical ionospheric hole produced by a man‐made explosion. Key Points: The 18 November 2023 Starship flight and explosions generated large‐scale multi‐oscillation supersonic conic waves in the ionosphereThe cone angle of the V‐shaped ionospheric disturbances corresponds to the Mach angle of shock waves propagating in the lower ionosphereThe shock waves from the Starship explosion caused a depletion in total electron content (TEC) [ABSTRACT FROM AUTHOR]

In the work i tis investigated the intense coudition of lirnear viscoelastic (Maxwell type) thread at a cross-section blow by wedge havig flat frontal part. I tis supposed that a defeectiv part of the thread fits the surface of the blowing body. The problem is solved at a subsonic mode with the condition on threads break. In many fields of modern industry, materials having elastic as well as viscous properties are used. The given thesis is dedicated to one of such questions, i.e. there is studied a behavior of a viscous-elastic fiber (Maxwells model) at a normal blow on it with a rigid wedge and a wedge with a flat front part and constant velocity. [ABSTRACT FROM AUTHOR]

A self-aligned focusing schlieren (SAFS) system combines the field of view of a conventional schlieren system with the defocus blur of a focusing schlieren system away from the object plane. It can be assembled in a compact form, measuring 1.2 m (4 ft) in length in the described case. The depth of field is sufficiently shallow to distinguish specific spanwise features in a supersonic flow field within a 76.2 mm (3 in) wide test section. As a result, the boundary-layer perturbations on windows and window-material defects and surface imperfections are blurred. Analytical forms are derived for depth of field and vignetting of the SAFS system. A laser spark velocity measurement in Mach 2 flow is performed by tracking the blast wave of a laser spark using 500 kHz SAFS imaging with a 200 ns optical pulse width. The flow Mach number and stagnation temperature are measured by comparing the blast-wave dynamics to an analytical solution. Additionally, schlieren image velocimetry is performed by analyzing natural flow perturbations in 500 kHz SAFS images using a self-correlation method. Comparing the spectra of gas density perturbations from the core flow and a near-wall region reveals a significant difference, with high-frequency prevalence at the boundary-layer location. [ABSTRACT FROM AUTHOR]

The oblique detonation waves have been studied extensively, but the wave characteristics influenced by the geometrical restriction have not been fully addressed. In this study, we examine the wave stability and the thrust performance in a contraction-expansion channel with varying velocity. A supersonic stoichiometric inflow of hydrogen-oxygen inflow is established in the computational domain, and the compressible reactive Navier-Stokes equations are solved using a comprehensive chemical model. As the velocity increases, the detonation wave inside the channel exhibits two successive unsteady states: the half normal detonation wave (half NDW) and the re-ignition oblique detonation wave (re-ignition ODW). In the half NDW state, the upper part of the wave surface is a basically stable NDW, while the lower part oscillates regularly as the reaction front. In the re-ignition ODW state, the explosion of the reaction front and the retreat of the detonation surface occur in proper order. Furthermore, the thrust associated with these newly discovered oscillation wave systems exhibits unstable behavior, with an important observation that it does not consistently decrease with increasing velocity. Notably, there is a significant increase in thrust during the transition from the half NDW state to the ODW state, as well as when the position of the oblique detonation wave shifts downstream. [Display omitted] • The oblique detonation wave structure and thrust performance in a contraction–expansion channel are revealed. • With increasing velocity, thermal choking unstarting, half normal detonation, and oblique detonation states emerge. • Two unsteady wave states were identified for the first time, and the primary flow characteristics were discerned. • Throughout the entire super-detonation process, the effect of V in / V cj on thrust is non-monotonic decreasing. [ABSTRACT FROM AUTHOR]

Supersonic and hypersonic flows have gained considerable attention in the aerospace industry in recent years. Flow control is crucial for refining the quality of these high-speed flows and improving the performance and safety of fast aircraft. This paper discusses the distinctive characteristics of supersonic flows compared to low-speed flows, including phenomena such as boundary layer transition, shock waves, and sonic boom. These traits give rise to significant challenges related to drag, noise, and heat. Therefore, a review of several active and passive control strategies is provided, highlighting their significant advancements in flow transitions, reducing drag, minimizing noise, and managing heat. Furthermore, we provide a comprehensive analysis of various research methodologies used in the application of flow control engineering, including wind tunnel testing, flight testing, and computational fluid dynamics (CFD). This work gives an overview of the present state of flow control research and offers insights into potential future advancements.

Low-fidelity analytic and computational wave drag prediction methods assume linear aerodynamics and small perturbations to the flow. Hence, these methods are typically accurate for only very slender geometries. The present work assesses the accuracy of these methods relative to high-fidelity Euler, compressible computational-fluid-dynamics solutions for a set of axisymmetric geometries with varying radius-to-length ratios (R / L) . Grid-resolution studies are included for all computational results to ensure grid-resolved results. Results show that the low-fidelity analytic and computational methods match the Euler CFD predictions to around a single drag count (∼ 1.0 × 10 − 4 ) for geometries with R / L ≤ 0.05 and Mach numbers from 1.1 to 2.0. The difference in predicted wave drag rapidly increases, to over 30 drag counts in some cases, for geometries approaching R / L ≈ 0.1 , indicating that the slender-body assumption of linear supersonic theory is violated for larger radius-to-length ratios. All three methods considered predict that the wave drag coefficient is nearly independent of Mach number for the geometries included in this study. Results of the study can be used to validate other numerical models and estimate the error in low-fidelity analytic and computational methods for predicting wave drag of axisymmetric geometries, depending on radius-to-length ratios. [ABSTRACT FROM AUTHOR]

Abstract On 18 November 2023, SpaceX launched the Starship, the tallest and the most powerful rocket ever built. The Super Heavy engine separated from the Starship spacecraft and exploded at 90 km of altitude, while the main core Starship continued to rise up to 149 km and exploded after ∼8 min of flight. In this work, we used data from ground‐based GNSS receivers and we analyzed total electron content (TEC) response to the Starship flight and the two explosions. For the first time, we observed large‐distance northward propagation of intensive 2,000 km V‐shaped ionospheric disturbances from the rocket trajectory. The observed perturbations, most likely, represent shock waves propagating with the cone angle of ∼14° on the North and ∼7° on the South against the flight track that corresponds to the Mach angle of the shock waves in the lower atmosphere. The Starship explosion also produced a non‐chemical depletion in the ionospheric TEC.

Rotating nozzle profile enables the supersonic variable Mach number wind tunnel to continuously adjust the Mach number of the experimental area during a single operation，which is important for studying the aerodynam ic problems in the maneuvering process of the aircraft and the starting process of the inlet. In the process of controlling the nozzle profile rotation，whether the flow field parameters can change linearly is an important criterion to measure the performance of supersonic variable Mach number wind tunnel. Therefore，the linear variation law of flow field parameters in the experimental area of variable Mach number wind tunnel is analyzed in this paper，and a numerical simulation model is established using the spring smoothing dynamic grid technique to verify whether the flow field parameters in the experimental area meet the linear variation law when the nozzle is located at the corresponding position of Mach number 3.041~3.215. The results show that the linear change of the flow field parameters in the wind tunnel experimental area is achieved by controlling the rotation of the nozzle profile，and the dynam ic calculation results are in good agreement with the expected linear change law of the flow field parameters in the experimental area. In the linear change process of flow field parameters with different accelerations，the deviation between the average parameters of the experimental area and the expected parameters at each time is less than 0.13%.

The process from test starting to supersonic flow state established in an outlet-opening wind tunnel，i. e. the starting problem of supersonic flow，has become a recognized challenge. In order to establish the feasible flow field starting method，and lay the foundation for the use of outlet-opening wind tunnel，three-dimensional numerical simulation research is carried out，which is based on a certain wind tunnel，and taken the panel blade of supersonic compressor as the research object. The failure reason of supersonic flow field starting on experimental condition is analyzed，and three flow field starting schemes are established. The results show that the starting failure reason is the formation of a strong normal shock wave at the leading edge of the cascade. Increasing the inlet total pressure alone cannot establish the supersonic flow. The supersonic cascade flow field can be started by increasing the overflow gap width on lower wall side，but the number of effective cascade passages is decreased and the boundary layer thickness is increased. Maintaining the width of overflow gap on the upper and lower walls at more than 1.0 pitch，the setting upper and lower supersonic walls in front of cascade and conducting suction can effectively start the supersonic flow field. The maximum fluctuation of outlet Mach number is 0.01，the maximum fluctuation of outlet flow angle is 0.09°. The periodicity can meet the experimental requirements.

This work presents the results of two-dimensional numerical simulation of processes, arising from the interaction of an incoming flow of a well-mixed stoichiometric hydrogen-air mixture with a wedge in a bounded region at high Mach velocities using Navier-Stokes type equations taking into account chemical interactions. The Mach number of the incoming flow from 9 to 12 were considered. Various wave configurations were obtained depending on the velocity of the incoming flow. For the flow Mach number of 9, the presence of a detonation wave perpendicular to the flow was obtained. For high flow velocities, another similar wave configuration of the same type was obtained. • A simulation of a flow interaction with a wedge at high Mach speeds was carried out • A well-mixed stoichiometric hydrogen-air mixture was considered • Hong's kinetic mechanism was considered • Two wave configurations were obtained depending on the velocity of the flow • For Mach number equal 9 the detonation wave perpendicular to the flow was obtained [ABSTRACT FROM AUTHOR]

With growing interest in sustainable civil supersonic and hypersonic aviation, there is a need to model the combustion of alternative, sustainable jet fuels. This work presents numerical simulations of several related phenomena, including laminar flames, ignition, and spray flames. Two conventional jet fuels, Jet A and JP-5, and two alternative jet fuels, C1 and C5, are targeted. The laminar burning velocities of these fuels are predicted using skeletal and detailed reaction mechanisms. The ignition delay times are predicted in the context of dual-mode ramjet engines. Large Eddy Simulations (LES) of spray combustion in an aeroengine are carried out to investigate how the different thermodynamic and chemical properties of alternative fuels lead to different emergent behavior. A novel set of thermodynamic correlations are developed for the spray model. The laminar burning velocity predictions are normalized by heat of combustion to reveal a more distinct fuel trend, with C1 burning slowest and C5 fastest. The ignition results highlight the contributions of the Negative Temperature Coefficient (NTC) effect, equivalence ratio, and hydrogen enrichment in determining ignition time scales in dual-mode ramjet engines. The spray results reveal that the volatile alternative jet fuels have short penetration depths and that the flame of the most chemically divergent fuel (C1) stabilizes relatively close to the spray. [ABSTRACT FROM AUTHOR]

An experimental study on the supersonic flow of Mach number 1.81 Over the rectangular cavity of L/D = 3 with different aft wall angles such as Baseline case 90° and aft wall edge ramp of 30°. Measurements of acoustic oscillations were carried out on the wall and floor of the cavity using unsteady pressure transducers. Schlieren flow visualization indicates the presence of a shear layer and the other shock features that are associated with the cavity flow field. In order to obtain the experimental results, data analysis was performed in the unsteady data Using statistical analytical techniques i.e., Fast Fourier Transform, Short Time Fourier Transform, and Coherence. After careful examination of the power spectra of the cavities, it is concluded that as the aft wall angle decreases by introducing the aft wall ramp, the amplitude decreases. It is clearly observed that the complete escape of vortices in Low angled cavity. From the correlation plot, the existence of the acoustic wave inside the cavity for all the cases is deduced. The maximum OASPL and static pressure values are observed at the aft wall. These results, the strength of the self-sustained oscillations and acoustic disturbance and mode switching were observed in the higher angled cavity 90°. Also, the mode switching co-existence was reduced in the aft wall edge ramp angle of 30°. [ABSTRACT FROM AUTHOR]

This present study describes the aerodynamic characteristics of supersonic flow over biconvex and double wedge airfoils using a finite volume method-based commercial CFD code Ansys Fluent. A steady-state RANS approach is used with SST k-ω viscous modeling. A series of simulations are conducted to analyze the characteristics of shock and expansion waves formed around the airfoils for Mach numbers ranging from 1.4 to 3.4 with varying angles of attack (α) from 0° to 20°. It is observed that the lift and drag coefficients both increase with the angle of attack for a fixed Mach number and decrease with the Mach number for a fixed angle of attack. Double wedge airfoil generates about 5% more lift at a low Mach number and 1% more lift at a higher Mach number compared to the biconvex airfoil. However, the biconvex airfoil generates lesser drag than the double wedge airfoil. The maximum value of the pressure coefficient (Cp) is found to be 1.7 for biconvex airfoil and 1.4 for double wedge airfoil. The maximum value of the lift-to-drag ratio for biconvex airfoil is 7.63, occurs at 1.4 Mach number and 3.46° angle of attack, whereas the value for double wedge airfoil is 5.19 at the same Mach number with 4.47° angle of attack, which suggest that biconvex airfoil has a higher lift-to-drag ratio and gives a better aerodynamics performance. The shock waves start to detach after an angle of attack of 5° and the shock wave is fully detached at a 15° angle of attack for biconvex airfoil for Mach number of 1.4. For the same Mach number, the double wedge airfoil, the shock wave starts to form the same as the biconvex airfoil but the waves are fully detached at a lower angle of attack of 10°. With the increasing Mach number, the shock waves remain attached to the airfoil. [ABSTRACT FROM AUTHOR]

A self-aligned focusing schlieren (SAFS) system combines the field of view of a conventional schlieren system with the defocus blur of a focusing schlieren system away from the object plane. It can be assembled in a compact form, measuring 1.2 m (4 ft) in length in the described case. The depth of field is sufficiently shallow to distinguish specific spanwise features in a supersonic flow field within a 76.2 mm (3 in) wide test section. As a result, the boundary-layer perturbations on windows and window-material defects and surface imperfections are blurred. Analytical forms are derived for depth of field and vignetting of the SAFS system. A laser spark velocity measurement in Mach 2 flow is performed by tracking the blast wave of a laser spark using 500 kHz SAFS imaging with a 200 ns optical pulse width. The flow Mach number and stagnation temperature are measured by comparing the blast-wave dynamics to an analytical solution. Additionally, schlieren image velocimetry is performed by analyzing natural flow perturbations in 500 kHz SAFS images using a self-correlation method. Comparing the spectra of gas density perturbations from the core flow and a near-wall region reveals a significant difference, with high-frequency prevalence at the boundary-layer location.

Axisymmetric direct numerical simulation (DNS) has been carried out to predict supersonic base flow behavior. Substantially fine grid has been used to perform calculations for the flow with Reynolds number up to 106. Optimal grid resolution was established through test calculations for affordable run time and solution convergence determined by the vorticity value. Numerical scheme provides fourth-order approximation for dissipative, fifth-order for convective and second-order for unsteady terms of conservation equations. Reynolds Averaged Navier-Stokes (RANS) approach has been employed to obtain input flow profiles for DNS calculations. Series of calculations have been carried out for Mach number 1.5 with Reynolds numbers 104, 105, 106 and for Mach number 2.46 with Reynolds number 1.65×106. It has been found that local base pressure coefficient calculated by DNS is a bit overestimated in a zone close to symmetry axis in comparison with experiment while integrated base drag coefficient shows good agreement with experimental data and noticeably better than one obtained by RANS approach.

Partially or completely reusing launch vehicles significantly reduces expenses for manufacturing and testing in space missions. Thus, Reusable Launch Vehicles (RLV) can be a solution to achieving technological and financial goals associated with space programs. Scramjet engines have a higher specific impulse than rocket engines and the capability to operate efficiently at extremely high speeds. This paper reviews the requirements in reusable launch vehicle technologies and how the scramjet engines can be used with their advantages and limitations, and possible solutions to meet future requirements of reusable launch systems. The review covers different aspects of the launch requirements of a reusable launch vehicle, the applicability of scramjet in the RLV, and the key parameters to enhance the performance of the scramjet engine. With the possibility of hydrogen as fuel, it is inferred that the scramjet is a plausible candidate for the successful, sustainable, greener, and cost-effective operation of the RLV. The combination of Scramjet-RLV holds the promise of significantly bringing down the cost of space missions. • Requirements in reusable launch vehicle technologies are reviewed. • Capabilities of scramjets to meet reusable launch system requirements are assessed. • With hydrogen fuel, scramjet can enable sustainable and cost-effective RLV operation. • Scramjet-RLV combine promises significant cost reduction in space missions. [ABSTRACT FROM AUTHOR]

Tracer pathline statistics have rarely, if at all, been used in the investigation of compressible subsonic, transonic and supersonic flows past a circular cylinder. Here, we report pathline histogram, conditional histogram and zonal histogram data, focusing on the characteristics of particle instantaneous flow reversal and their dependencies on Mach number and tracer release position. The far-upstream Mach number and the diameter-based Reynolds number used for these three flow regimes are $ ({0.2, 10000}), ({0.9, 3900}) $ (0.2 , 10000) , (0.9 , 3900) and $ ({1.2, 10000}) $ (1.2 , 10000) , respectively. The subsonic surface-release histogram profiles display unexpected persistent oscillations at a frequency $ 50 $ 50 times of the periodic vortex shedding frequency, reflecting the effect of a very small magnitude zig-zagging type of particle motion in a region very close to the wall and slightly behind the mean flow separation location. No similar high-frequency oscillation is observed in the corresponding transonic and supersonic histograms likely because, at $ \textrm{Ma} \ge 0.9 $ Ma ≥ 0.9 , periodic vortex shedding ceases to exist and the near-wake becomes a quasi-laminar recirculation zone bounded by a pair of converging slip-layers with a neck opening to the far-wake. At the instant of $ 5 $ 5 subsonic vortex-shedding periods after being released from the cylinder surface, there are $ 1.67\% $ 1.67 % , $ 35.2\% $ 35.2 % and $ 39.7\% $ 39.7 % particles that experience flow reversal in the subsonic, transonic and supersonic flows, respectively. Surprisingly, in the present transonic flow, growth of the turbulent far-wake behind the pair of lambda-shocks is constrained over a considerable streamwise distance. [ABSTRACT FROM AUTHOR]

This study investigates the thermal stressing characteristics of hydrocarbon fuels to develop thermally stable fuels for supersonic vehicle applications. The thermal stability characteristics of two multicomponent hydrocarbon fuels, namely F-1 and F-2 fuels, under subcritical and supercritical conditions, are examined for a wide range of temperatures (30–500 °C) and pressures (15–50 bar) using a semi-batch reactor. Various analysis techniques, such as UV-Vis spectrophotometry, ASTM D86 distillation, gas chromatography, scanning electron microscopy (SEM), etc., are used to characterize the feed and product properties. The study showed that the fuels are reasonably stable up to around 400 °C and above 40 bar pressure. The stability of the fuels is relatively less under subcritical pressure at elevated temperatures. The amount of gum formation increased with the increase of both temperature and pressure parameters. Further, the SEM analysis showed oval-shaped and ribbon-like structures in the solid deposits. The present investigation may be useful in developing an appropriate fuel for regenerative cooling in supersonic applications. [ABSTRACT FROM AUTHOR]

In the present study, the mixing characteristics of the Mach 1.86 jet from a convergent-divergent nozzle of square cross-section from inlet to exit were investigated numerically. The jet studied were uncontrolled and controlled with limiting tab of rectangular and circular cross-sections. The nozzle pressure ratio is varied from 4 to 6 with a step size of one, leading to the confinement of the present study for the over-expansion and near correct expansion levels of the jet. It is observed that the core length, which is a direct indication of the extent of jet mixing, reduced significantly for the jet controlled with limiting tab as compared to that with the uncontrolled counterpart. Also, among the circular and rectangular limiting tabs, the tab with rectangular cross-section caused maximum core length reduction. The jet spread and the waves prevailing in the jet field were studied and visualized using pressure profiles and Mach contours, respectively. The present work may be used in understanding and solving complex phenomena occurring in the fields of aeroacoustics noise suppression, reduction of base heating of launch vehicles, supersonic combustion, mitigation of infrared radiations due to chimney smoke, etc., where jet mixing is highly desirable. [ABSTRACT FROM AUTHOR]

Aerodynamic interference is one of the phenomena that occurs when two flying objects pass near each other. In this case, the change in the pressure distribution on objects passing near each other causes a change in the aerodynamic forces. In this research, the numerical investigation of the flow field between two narrow bodies at a close distance and the pressure changes along them at supersonic speeds is done. The simulation was done in two-dimensional and three-dimensional form, and the k-omega sst was used to model the flow turbulence. The two slender bodies were placed next to each other at speeds of Mach 1.5, 3, and 4.5 and at intervals of 2, 3, and 4 times the diameter of the body. they were examined in two modes with relative motion and without relative motion. The results show that as the free flow velocity increases, the shock wave deforms from bow to oblique and is reflected between the two bodies. Also, as the flow Mach number decreases and the distance between two slender bodies increases, the reflected shock wave becomes weaker. At Mach 1.5 the shock wave reflection between two bodies is not very noticeable, while at Mach 4.5 the shock wave is well reflected between two bodies. In the state of relative motion of two slender bodies, separation occurs at the shock wave reflection points and the number of shock wave reflections decreases compared to the state without relative movement.

Low-fidelity analytic and computational wave drag prediction methods assume linear aerodynamics and small perturbations to the flow. Hence, these methods are typically accurate for only very slender geometries. The present work assesses the accuracy of these methods relative to high-fidelity Euler, compressible computational-fluid-dynamics solutions for a set of axisymmetric geometries with varying radius-to-length ratios (R/L). Grid-resolution studies are included for all computational results to ensure grid-resolved results. Results show that the low-fidelity analytic and computational methods match the Euler CFD predictions to around a single drag count ( ∼1.0×10−4) for geometries with R/L≤0.05 and Mach numbers from 1.1 to 2.0. The difference in predicted wave drag rapidly increases, to over 30 drag counts in some cases, for geometries approaching R/L≈0.1, indicating that the slender-body assumption of linear supersonic theory is violated for larger radius-to-length ratios. All three methods considered predict that the wave drag coefficient is nearly independent of Mach number for the geometries included in this study. Results of the study can be used to validate other numerical models and estimate the error in low-fidelity analytic and computational methods for predicting wave drag of axisymmetric geometries, depending on radius-to-length ratios.

Measurement while drilling (MWD) plays an important role in controlling well trajectory, improving drilling efficiency and saving drilling cost. In recent years, a large number of high-temperature wells have appeared, which makes the traditional MWD instruments no longer applicable. So far, no active cooling method has been successfully applied to MWD instruments for high-temperature oil and gas wells. This paper presents a new active cooling method for air drilling by utilizing the throttling and cooling of drilling gas, and the cooling system is designed. An axisymmetric numerical model of internal fluid is established, and different drilling gas flow channel design methods are discussed. The research results indicate that the wall temperature of the cooling section is much higher than the center fluid temperature, but lower than the ambient temperature, and can be used for active cooling of downhole drilling instruments. Compared with other design methods, the supersonic nozzle designed using the Bezier method has better parallelism of the gas in cooling section. The diffusion capability of straight wall diffuser is better than that of curved wall diffuser, and it can withstand greater back pressure. For straight wall diffuser, increasing the throat length and reducing the contraction and expansion angle can improve the diffusion capacity, but they also lead to a longer diffuser.

Axisymmetric direct numerical simulation (DNS) has been carried out to predict supersonic base flow behavior. Substantially fine grid has been used to perform calculations for the flow with Reynolds number up to 106. Optimal grid resolution was established through test calculations for affordable run time and solution convergence determined by the vorticity value. Numerical scheme provides fourth-order approximation for dissipative, fifth-order for convective and second-order for unsteady terms of conservation equations. Reynolds Averaged Navier-Stokes (RANS) approach has been employed to obtain input flow profiles for DNS calculations. Series of calculations have been carried out for Mach number 1.5 with Reynolds numbers 104, 105, 106 and for Mach number 2.46 with Reynolds number 1.65·106. It has been found that local base pressure coefficient calculated by DNS is a bit overestimated in a zone close to symmetry axis in comparison with experiment while integrated base drag coefficient shows good agreement with experimental data and noticeably better than one obtained by RANS approach. [ABSTRACT FROM AUTHOR]

The complex fluid-dynamic instabilities and shock waves occurring along the surface of a two-dimensional wedge at high values of the Mach number are studied here through numerical solution of the governing equations. Moreover, a regression model is implemented to determine the pressure distribution for various Mach numbers and angles of incidence. The Mach number spans the interval from 1.5 to 12. The wedge angles (θ) are from 5° to 25°. The pressure ratio (P2/P1) is reported at various locations (x/L) along the 2D wedge. The results of the numerical simulations are compared with the regression model showing good agreement. [ABSTRACT FROM AUTHOR]

In computational fluid dynamics of compressible fluid flow, the simple low‐dissipation advection upstream (SLAU) scheme formulated with multidimensional velocity components (normal and parallel to a cell interface) is a widely employed all‐speed scheme. As a variant of SLAU, the mSLAU scheme, which adopts only a velocity component normal to the cell interface instead of multidimensional velocity components, is used for rotorcraft calculations. However, although mSLAU has been claimed to be empirically stable, it has been pointed out that using only the cell‐interface‐normal velocity component instead of the multidimensional velocity components causes numerical instability. Therefore, to clarify the roles of the multidimensional velocity components for computational stability, we solved some benchmark problems associated with using SLAU or mSLAU. We discovered that the multidimensional velocity components contributed to stability against poor‐quality grids by isotropically producing enough amount of numerical dissipation, especially in low‐subsonic and supersonic flows. Although mSLAU could practically treat moderate Mach number flows (approximately 0.1 < M < 1.0) when coupled with the minmod limiter, using only the cell‐interface‐normal velocity component can deteriorate convergence of calculations and lead to susceptibility in the grid geometry. [ABSTRACT FROM AUTHOR]

The article discusses the current state of development of supersonic transport, analyses the main limitation of the use of gas turbine engines when flying at high supersonic flight speeds, proposes a method for expanding the range of GTE limitations in these flight modes of the aircraft, using cold forcing of the engine, conducts an analytical study of the effectiveness of this approach, proposes one of the possible devices for the implementation of cold forcing, the efficiency of an engine with cold forcing is shown, which may be of significant interest for the application of cold forcing methods in gas turbine engines of supersonic transport.

In this paper, one of the remaining questions for ejector design, “Does the constant rate of momentum change (CRMC) ejector provide better performance than the constant pressure mixing (CPM) ejector under identical ejector area ratio, ejector length, and operating conditions?”, has been answered. Two steam ejector designs operating with the various boiler and evaporator temperatures were simulated by the SST k-omega turbulence model. The present computational fluid dynamics (CFD) model produced a reasonable agreement with our previous published experimental data. The upstream operating conditions were simultaneously considered to assess the better performance ejector design with the help of the primary expansion coefficient. The predicted results revealed that the CRMC ejector showed an advantage in entrainment ratio and a disadvantage in critical condenser pressure. However, the ejector efficiency comparison confirmed that the CRMC ejector design provided better performance than the CPM ejector design for a primary expansion coefficient greater than unity. The maximum percentage improvement of ejector efficiency was 32.418%.

Dynamic homogenization of heterogeneous materials has received a lot of attention recently, due to the fact that it can avoid expensive direct simulations with high space-time resolution. In this paper, the applicability of effective homogenous media on some peculiar dynamic phenomena is discussed. Longitudinal wave velocity has been believed to be the upper limit of mechanical signal speeds in a homogenous elastic solid according to the classical elasticity theory. However, for heterogeneous materials, we demonstrate numerically and theoretically that if the wave impedances of the constituent phases are not equal, there must be the supersonic phenomenon such that some wave may propagate faster than the longitudinal wave in the effective homogenous medium. Especially, if two materials with the same longitudinal wave velocity are used to compose a heterogeneous medium, the effective longitudinal wave speed will almost definitely be lower than its local counterpart. We also investigate the evolution of energy distribution over time when the wave propagates through heterogeneous media with various gradients, which indicates, interestingly, that the smoother the change in wave impedances of the heterogeneous medium, the more energy in "before-wave" region and the less energy in "after-wave" region. The corresponding dispersion curves and group speed curves also show dramatic change for heterogeneous media with different smoothness, which can partially explain but not quantitively predict these phenomena. We attempt to develop an effective medium model to reproduce the energy evolution but fail. All these findings pose a big challenge to the proper establishment of dynamic homogenization models. [ABSTRACT FROM AUTHOR]

In this study, a duct is considered and special attention is paid to a passive method for the control of the base pressure relying on the use of a cavity with a variable aspect ratio. The Mach number considered is 1.8, and the area ratio of the duct is 2.56. In particular, two cavities are examined, their sizes being 3:3 and 6:3. The used L/D spans the interval 1--10 while the NPRs (nozzle pressure ratio) range from 2 to 9. The results show that the control becomes effective once the nozzles are correctly expanded or under-expanded. The pressure contours at different NPR and L/D are presented. It is shown that the NPR and cavity location strongly influence the base pressure. The NPR, Mach number, and cavity aspect ratio have a strong effect on the base pressure in the wake region. [ABSTRACT FROM AUTHOR]

The back pressure at the outlet of supersonic cyclone separator usually causes aerodynamic shock wave in Laval nozzle, and the interaction of supersonic flow, shock wave and phase transition is very complicated. In order to explore the supersonic flow properties of wet gas under the action of shock wave, study was performed on the methane-water vapor mixtures. Then, a mathematical model for the supersonic condensation and flow of wet natural gas was established. Meanwhile, the effect of the outlet back pressure and inlet pressure on the supersonic condensation property of wet natural gas was analyzed to evaluate the performance of the supersonic cyclone separator under the action of shock waves. The simulation results show that the shock waves can cause the rapid increase in pressure and temperature of water vapor in a short time, leading to the condensed droplets re-evaporate and the liquid mass fraction at outlet reduces sharply. Besides, as the inlet pressure increases, the location of nucleation approaches the throat of Laval nozzle gradually, and the limit nucleation rate and the droplet number decrease accordingly. However, the liquid mass fraction increases due to the increasing of growth rate and sizes of droplets. Moreover, the nozzle loses its function of liquefaction when the back pressure is higher than 25 kPa (i.e., the pressure ratio is higher than 0.58), and then the nozzle cannot operate normally. Therefore, it is necessary to control the reasonable pressure ratio to ensure the dehydration efficiency of the supersonic separator with back pressure.

Ignition and relighting in supersonic flows is an important challenge for the design of hypersonic propulsion systems. Supersonic compressible flows of interest exhibit much larger local variations in velocity, shear, and thermodynamic state than their incompressible counterparts. Thus, it is of interest to study the relationship between ignition kernel evolution, the initial spark location, and the kernel's subsequent flow state history. We leverage the flexibility of a laser plasma ignition system to systematically explore a large set of spark locations on the symmetry plane of an ethylene jet in supersonic crossflow setup. CH* measurements are used to visualize chemically active regions and results are correlated with flow field properties derived from Mie-scattering data of the non-reacting flow field. Our study describes the laser plasma properties in detail and scrutinizes the effect of turbulent mixing and flow dilatation on ignition kernels. Finally, the results yield general guidelines for favorable ignition locations in the engineering design of chemically reactive compressible flows. [ABSTRACT FROM AUTHOR]

The article discusses the current state of development of supersonic transport, analyses the main limitation of the use of gas turbine engines when flying at high supersonic flight speeds, proposes a method for expanding the range of GTE limitations in these flight modes of the aircraft, using cold forcing of the engine, conducts an analytical study of the effectiveness of this approach, proposes one of the possible devices for the implementation of cold forcing, the efficiency of an engine with cold forcing is shown, which may be of significant interest for the application of cold forcing methods in gas turbine engines of supersonic transport. [ABSTRACT FROM AUTHOR]

Aero-acoustic characteristics of a high-speed jet with elliptical ring tabs fitted at the aft of the nozzle in diametrically opposite to the exit is experimentally investigated at a Mach Number of 1.8. The Main focus is to examine the effects of the extent of elliptical ring tabs, its position in the mixing layer and also to study the characteristics of supersonic nozzle with and without elliptical ring tabs. The elliptical ring tabs are found to produce vortices, thereby increasing the surface area of the jet, particularly in the close vicinity of the nozzle, which increases the mixing and reduces the potential core length. The overall noise levels registered along the jet edge immediately downstream of the nozzles are higher, but further downstream they are reduced in comparison with the plain baseline nozzle. A pair of dimensionally optimized elliptical ring tab placed diametrically opposite at the exit of the nozzle with blockage area of 5% without effecting the loss of thrust. The mixing enhancement caused by the tab will be studied at nozzle pressure ratio 5.75. The reduction in Sound Pressure Level (SPL) in dB is compared to that of plain nozzle. [ABSTRACT FROM AUTHOR]

Energy deposition for flow and flight control has received significant interest in the past several decades due to its potential application to high-speed flow and flight control. This paper reviews recent progress and recommends future research. [ABSTRACT FROM AUTHOR]

• The flutter mechanism of the "diamond-back" folding wing under supersonic inflow is investigated. • The characteristics of the flutter boundary of the "diamond-back" folding wing with Mach number and swept-back angle are studied. • The flutter evolution mechanism is analyzed by using the characteristics of aerodynamic and flow field, instability characteristics of system dynamics, and characteristics of structural natural mode. • The "jump" phenomenon on the flutter boundary is found and analyzed in detail. Folding wings with adjustable geometry offer a viable solution for vehicles to adapt to complex and variable flight environments, while enhancing aerodynamic performance. This paper investigates the flutter mechanism and characteristics of the "diamond-back" folding wing under supersonic inflow by a ROM-based aeroelastic method. The evolution mechanism is analyzed by the characteristics of aerodynamic and flow field, system dynamics, and structural natural mode. The findings reveal that, firstly, as the Mach number increases, the flutter critical speed monotonically increases, while the flutter critical frequency exhibits fluctuations within a specific range. Secondly, the study reveals a pivotal transformation in the modal shapes at a 35° swept-back angle that reshapes the system dynamics, transitioning from second and third-order coupled instability to first and second-order coupled instability. Consequently, this nonlinear change significantly amplifies the flutter critical speed and reduces the flutter critical frequency. Lastly, the combined effect of Mach number and swept-back angle significantly influences the local flutter boundary. Under low supersonic speeds and large swept-back angles, a bow shock forms in the local flow field region, resulting in notable changes in pressure distribution. As the Mach number increases, the local flutter boundary exhibits a nonlinear trend of first increasing, then decreasing, and then increasing again. Conversely, as the swept-back angle increases, it demonstrates a nonlinear trend of first decreasing, then increasing. These insights provide valuable guidance for the aerodynamic and structural refinement of "diamond-back" folding wings, catering to sophisticated flight dynamics and environmental adaptability. [ABSTRACT FROM AUTHOR]