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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]

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.

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.

• A gust response alleviation approach is investigated for the supersonic wing aeroelastic systems. • Effect of the nonlinear energy sink (NES) on the aeroelastic gust response is studied. • Influence of NES parameters on the dynamic response amplitude is discussed. • The optimal relative distance for the installation of NES is found. • Influence of nonlinear energy sink on the targeted transfer of vibration energy is analyzed. In order to reduce the gust response level of the aircraft wing aeroelastic systems in supersonic flows, the nonlinear energy sink (NES) is introduced into the aeroelastic system, and the effect of the nonlinear energy sink on the aeroelastic gust response is studied. The gust response dynamic model without and with NES is established by Lagrange equation, considering the cubic nonlinear characteristics of both plunge and pitch stiffness. The piston theory model as well as the '1-cos' model are employed to describe the unsteady aerodynamic load and the gust velocity, respectively. Besides, the comparison study of the gust response amplitude and energy characteristics of the wing aeroelastic systems with and without NES are analyzed. For systems with NES, the amplitude of gust dynamic response decreases rapidly, while vibration can be transmitted to the NES mechanism and causing its vibration. From an energy perspective, about 72.15 % of the original vibration energy has been transferred into NES, and further consumed through its damping. The influence of NES parameters, such as frequency ratio, mass ratio, damping ratio, and installation position on gust response, is discussed. Numerical simulations demonstrate that: as the frequency ratio, damping ratio and mass ratio increase, the dynamic response amplitude of the aeroelastic system decays faster, and the energy transferred from the wing system also increases. In addition, an optimal relative distance for the installation of NES can be found with best gust alleviation effect. The influence of nonlinear energy sink on the targeted transfer of vibration energy in aeroelastic systems are analyzed under different working conditions. The NES shows great gust alleviation effect and good robustness under different flight speeds, gust intensities, and gust lengths. [ABSTRACT FROM AUTHOR]

• A concentration effectiveness of seeping gas film was visualized by PSP. • Attenuation law of film effectives downstream porous injector under different conditions was investigated. • The mixing behavior between seeping gas film and boundary layer was analyzed. Based on seepage flow through porous media, the seeping gas film cooling method is an effective means to protect large areas of optical windows and other hot components of hypersonic vehicles. Here, the pressure-sensitive paint (PSP) technique was applied and experiments were conducted on seeping gas film effectiveness downstream of a porous injector. A Mach 3 wind tunnel was used to explore the influence of different cooling gas blowing ratios, gas types, and incoming boundary-layer conditions on film effectiveness. Results show film effectiveness decreases monotonically along the downstream wall; however, with blowing ratio increases the film effectiveness increases linearly. Once the blowing ratio exceeds 0.4% film effectiveness ascends nonlinearly, indicating that the contribution to film effectiveness per unit mass flow-rate promotion gradually decreases. In contrast, the non-uniformity of the film coverage on the downstream wall in the spanwise direction becomes more significant at a higher blowing ratio. For the same blowing ratio, helium, with a low molecular weight has a higher film effectiveness compared to nitrogen and carbon dioxide. With blowing ratio less than 0.4 %, the film effectiveness downstream of the porous injector under laminar flow conditions is almost twice that of turbulent flow. However, with blowing ratio above 0.5 %, the growth rate of the film effectiveness decreases dramatically to 50 %. Preliminary analysis suggests that this is caused by the complete transition of the laminar boundary layer to turbulent flow after passing through the porous wall under high blowing ratios, where the mixing effect of turbulence is fully manifest. [ABSTRACT FROM AUTHOR]

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]

A comprehensive and systematic, computational investigation is presented on the effect of wall roughness on the supersonic flow over a circular cylinder with a Reynolds number of 500. Flow simulations are conducted using ANSYS Fluent. Wall roughness is modeled by a perturbation of the cylinder geometry with harmonic modes of varying amplitude and frequency. Validated smooth cylinder flow simulations for a range of Mach and Reynolds numbers with slip and no-slip wall serve as a reference. Roughness is shown to increase the effective diameter of the cylinder and the drag by displacing the outer flow along the peaks of the roughness elements. For lower frequencies, this effect is less pronounced than for higher roughness frequencies. While for smooth cylinders the vorticity is mostly generated by viscous shear forces, for rough cylinder the baroclinic vorticity generation is shown to be dominant and shown to determine the topology of the recirculating region. [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]

A vacuum supersonic ejector is an indispensable pneumatic device placed in nearly all industrial production lines. This device, also called a zero-secondary flow ejector, is characterized by the maximum entrained flow and the minimum secondary pressure. Numerical simulations were carried out by means of the CFD toolbox OpenFOAM v8 and its solver HiSA, which uses the AUSM+up upwind scheme. A single-factor analysis of eight parameters was performed to find how the ejector's performance was enhanced or decreased, while other parameters were fixed. Four parameters were subject to further analysis to find the geometry that improves the standalone performance of the ejector. The mixing chamber length is the parameter that most improves its performance; alone it leads to a 10% improvement. A multi-factor analysis, based on a fractional factorial design, is carried out with the four relevant parameters. Results indicate that the multi-factor analysis enhances the performance of the ejector by 10.4% and the mixing chamber length is the factor that most influences the improvement. Although a multi-factor design improves the performance, no significant relevance has been detected with respect to the mixing chamber length improvement alone. The improved performance of this device leads to a reduction in operating time and, as a consequence, results in significant energy savings. [ABSTRACT FROM AUTHOR]

Uncertainty quantification has proven to be an indispensable study for enhancing reliability and robustness of engineering systems in the early design phase. Single and multi-fidelity surrogate modelling methods have been used to replace the expensive high fidelity analyses which must be repeated many times for uncertainty quantification. However, since the number of analyses required to build an accurate surrogate model increases exponentially with the number of random input variables, most surrogate modelling methods suffer from the curse of dimensionality. As an alternative approach, the Low-Rank Approximation method can be applied to high-dimensional uncertainty quantification studies with a low computational cost, where the number of coefficients for building the surrogate model increases only linearly with the number of random input variables. In this study, the Low-Rank Approximation method is implemented for multi-fidelity applications with additive and multiplicative correction approaches to make the high-dimensional uncertainty quantification analysis more efficient and accurate. The developed uncertainty quantification methodology is tested on supersonic aircraft design problems and its predictions are compared with the results of single- and multi-fidelity Polynomial Chaos Expansion and Monte Carlo methods. For the same computational cost, the Low-Rank Approximation method outperformed both in surrogate modeling and uncertainty quantification cases for all the benchmarks and real-world engineering problems addressed in the present study. [ABSTRACT FROM AUTHOR]

Introduction: Non-alcoholic fatty liver disease (NAFLD) or fatty liver, is characterized by an excessive accumulation of fat in the liver, is a metabolic disorder with a worldwide prevalence close to 25%, with a spectrum of liver damage that covers the steatosis without fibrosis, steatohepatitis with variable fibrosis and cirrhosis or maximum degree of fibrosis, this fibrosis determines prognosis and outcomes in the disease. Objective: To evaluate the association between body mass index and the degree of liver fibrosis in patients diagnosed with fatty liver in a hepatology center in the city of Bogotá, Colombia. Patients and methods: A case-control study is carried out with patients diagnosed with fatty liver, who have undergone real-time elastography (Supersonic). Information was taken from patients diagnosed with fatty liver who met the inclusion criteria. Continuous variables were described using measures of central tendency and standard deviation. Categorical variables were described with numbers and percentages. A 95% confidence interval was considered statistically significant. Results: 361 patients were included, of which 95.2% (n=344) presented some degree of alteration (12% minimal fibrosis, 33% moderate fibrosis, 34% severe fibrosis and 16% cirrhosis) and only 5% showed a liver normal. Not having an adequate weight is related to severe fibrosis F3 OR 3.24 (1.03-10) and cirrhosis F4 OR 2.33 (2.33-42.99). No statistically significant differences were found between altered body mass index and any degree of fibrosis OR 2.74 (0.90-8.40). The presence of DM presents a 10-fold risk probability of ending in F4 cirrhosis, especially with poor disease control OR 5.16 (1.23-30.33). Conclusion: There is an association between abnormal body mass index and glycemic profile and the development of severe and advanced fibrosis. It is necessary in clinical practice, greater surveillance and evaluation of patients with fatty liver, in order to prevent the progression of fibrosis. [ABSTRACT FROM AUTHOR]

Separation of mixtures, gases, and phases is of great industrial importance. Various methods and equipment are utilized for the separation. One of the most novel and innovative as well as efficient methods is supersonic separation where acceleration of the flow mixture is accompanied by fast reduction in temperature and pressure. As a consequence, heavier components which have a higher condensation temperature will liquefy faster than lighter components. Swirling flow and centrifugal force will drive heavier condensed droplets toward the outer wall and separate them from lighter components in the gas phase. In general, supersonic separators consist of a converging-diverging nozzle and a cyclone separator. Due to no moving or rotating parts, supersonic separators have high reliability, long endurance, easy manufacturing, and less expensive operation and maintenance. In this paper, the one-dimensional design of a supersonic separator for separation and purifying of Methane from natural gas mixture is presented. Since separation process is in vicinity of saturation line, real gas model from NIST REFPROP software is used to estimate the fluid thermodynamic properties. For the design of cyclone separator and phase separations, instantaneous nucleation at the nozzle discharge is assumed and displacement of smallest droplet (as the worst case) is only considered. After calculating the supersonic separator geometry, performance sensitivity to changes in inlet temperature, pressure, and gas composition is studied, the effectiveness of these parameters is compared and rate of performance change is calculated. At next step, the simultaneous change of the inlet conditions is investigated and effect of each parameter on the effectiveness of others is studied. Finally, in order to optimize phase separation in various operating conditions, effects of swirler angle and swirl intensity on separation length are studied and optimum swirl intensity to gain the best possible separation performance with a fixed geometry separator for different inlet conditions is determined.

Performance of two axisymmetric air intakes are compared at conditions suitable for Mach number range from 2 to 8. First is the Busemann intake and second is the reversed isentropic nozzle. The isentropic nozzle is built by the method of characteristics. The contour of this nozzle is taken as a compression surface for the incoming flow. Performances of these two intakes are compared by comparison of both viscous and inviscid CFD calculations at Mach 6. Viscous flow calculations show that the total pressure recovery in compression section is 0.8316 in the Busemann intake and 0.869 in the reversed isentropic nozzle intake.

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We employ numerically implicit subgrid-scale modeling provided by the well-known streamlined upwind/Petrov–Galerkin stabilization for the finite element discretization of advection–diffusion problems in a Large Eddy Simulation (LES) approach. Whereas its original purpose was to provide sufficient algorithmic dissipation for a stable and convergent numerical method, more recently, it has been utilized as a subgrid-scale (SGS) model to account for the effect of small scales, unresolvable by the discretization. The freestream Mach number is 2.5, and direct comparison with a DNS database from our research group, as well as with experiments from the literature of adiabatic supersonic spatially turbulent boundary layers, is performed. Turbulent inflow conditions are generated via our dynamic rescaling–recycling approach, recently extended to high-speed flows. Focus is given to the assessment of the resolved Reynolds stresses. In addition, flow visualization is performed to obtain a much better insight into the physics of the flow. A weak compressibility effect is observed on thermal turbulent structures based on two-point correlations (IC vs. supersonic). The Reynolds analogy ( u ′ vs. t ′ ) approximately holds for the supersonic regime, but to a lesser extent than previously observed in incompressible (IC) turbulent boundary layers, where temperature was assumed as a passive scalar. A much longer power law behavior of the mean streamwise velocity is computed in the outer region when compared to the log law at Mach 2.5. Implicit LES has shown very good performance in Mach 2.5 adiabatic flat plates in terms of the mean flow (i.e., C f and U V D + ). iLES significantly overpredicts the peak values of u ′ , and consequently Reynolds shear stress peaks, in the buffer layer. However, excellent agreement between the turbulence intensities and Reynolds shear stresses is accomplished in the outer region by the present iLES with respect to the external DNS database at similar Reynolds numbers. [ABSTRACT FROM AUTHOR]

The present work provides a detailed study of a cloud seeding prototype rocket's nose cone drag reduction strategy by adopting the Computational Fluid Dynamics (CFD) simulation method for solving compressible flow problems. Two different nose cone aerodisk models are simulated using OpenFOAM software: two hemisphere aerodisk and two spherically blunted tangent ogive aerodisk. The results are compared against a baseline case of the blunt nose cone. The solver used is first validated against published experimental data. It is observed that both aerodisk models significantly reduce the drag coefficient of the blunt nose cone with two spherically blunted tangent ogive aerodisk providing the lowest drag coefficient. Detailed comparison of the flow visualization in terms of temperature distribution and shock wave formation is also reported to determine its influence on aerodynamic drag. The present study generally reveals that two spherically blunted tangent ogive aerodisk is preferred over two hemisphere aerodisk for achieving greater drag reductions and lower heat load to the cloud seeding prototype rocket. [ABSTRACT FROM AUTHOR]

This research examines the shockwave characteristics at a double ramp inlet and its performances at different angles of attacks (AoAs) with various Mach numbers. The k-omega (k-ω) shear stress transport (SST) turbulence model was used over a two-dimensional double ramp inlet at free-stream Mach numbers of 4, 5 and 6 with various AoAs in the range of -10 to 10°. The experimental data was used to validate the numerical results. Analysis of the internal shock structures showed that at a Mach number lower or greater than the intended Mach number, the shock waves' impact on the cowl inlet was considerably influenced. Moreover, the Mach number and AoA had major impact on overall pressure, kinetic energy and compression process efficiencies, whereby kinetic energy and compression process efficiencies improve when AoA is increased. An increase in AoA also shifts the shock wave inside the isolator, while decrease in AoA expels the shockwave from the isolator at the design Mach number. It can be concluded that the interaction of the shockwave with the cowl would be affected by various AoAs. [ABSTRACT FROM AUTHOR]

Prediction and limitation of inlet pressure are important for supersonic vehicles. Numerical simulation on high speed inlet flow is conducted, and pressure characteristics with different Mach numbers, attack angles, yaw angles and excess air coefficients is analyzed. Based on dimensionless and decoupled method, empirical formula for high speed inlet pressure is fitted and verified by flight results. The results show that, with the Mach number increasing and excess air coefficient decreasing, the pressure in inlet increases. Attack angle and yaw angle have little influence on inlet pressure in normal work envelope. The inlet pressure estimated by empirical formula is in good agreement with measured data in flight tests, which means that the formula established in this paper has a high precision.

In order to obtain the flow characteristics of sub-super-sonic mixing layer including velocity distribution, pressure distribution and development of mixing layer, experimental and numerical investigations were conducted. PIV technique was employed to measure the two-dimensional velocity distribution in the experiment while the standard k-ω turbulent considering the effect of compressibility was adopted to simulate the flow characteristic of mixing layer. The Mach number of subsonic stream and supersonic one was 0.11 and 1.32, respectively. The results show the flow of mixing layer is temporally transient. The interface between two streams lies initially as an approximately line segment; afterward, it becomes wrinkled and distorted; finally, it breaks up. The mixing layer develops linearly along streamwise direction in the time averaged velocity field with a growth rate of 0.135. The velocity and total pressure distributions in the mixing layer are self-similar.

Abstract— Supersonic intake buzz is a type of shock oscillation that could occur at any supersonic inlet of an aircraft and which can alter the mass flow and pressure at inlet. It may lead to combustion instability also. A lot of experiments have been performed to understand the buzz characteristics while visualizing the flow characteristics through flow visualization techniques. Also, there have been a great effort and research to understand the supersonic buzz phenomenon through numerical simulation with the help of computational software packages. There could be different parameters affecting the buzz. Some of them was discovered by the researchers all over the world. Intake buzz distortion get substantially affected by change in angle of attack. Decrease in throttle area was also found to produce prominent effect on buzz by increasing its dominant frequency. The suppression of buzz has become a topic of active research. The suppression of buzz can help in increasing the efficiency of supersonic inlet and efficient compression, so can result in benefit of increasing the economic output and reducing the fuel consumption. Intake performance was found to be increased with small cowl deflection and by the use of buzz predictor and margin controller. Buzz could be controlled by deflecting the air in an advantageous way and generating a weaker gradient at inlet lip. [ABSTRACT FROM AUTHOR]

Scramjet is an air breathing engine in which the combustion takes in a supersonic air stream by generating high pressure. Wedge shaped strut plays a major role in the scramjet engine to mix and compress air at high speeds, igniting the air fuel mixture and sustaining the flame. But the geometric profile of wedge in which the air gets blocked partially or completely is unanswered. Therefore, the effect of blockage on intake characteristics is investigated in this present study. Numerical investigations are carried out using ANSYS Fluent at Mach 5 to study the changes in static pressure, stagnation pressure, temperature and mass flow rate for different percentage of blockage which is varied from 10% to 45%. Unstart-free air intake for scramjet engine is ensured by keeping the percentage blockage of fuel injection struts between 10% and 35%. The Mach number at supersonic regime until 35% of blockage turns suddenly into subsonic conditions afterwards for further increase in blockage percentage. Moreover, as the blockage is further increased beyond 35%, bow shock is observed ahead of the cowl and the flow through the intake is subsonic, thereby indicating 'unstart' intake. Based on this geometry in air intake, smooth gas dynamic compression or ramming effect is obtained up to 35% of blockage. [ABSTRACT FROM AUTHOR]

An experimental study on the atomization characteristics of a liquid jet in supersonic crossflow is conducted to analyze fuel atomization in a high-speed air-breathing propulsion system. Phase Doppler Anemometry (PDA) is used to obtain the droplet size and cross-sectional shape. The influencing factors are investigated experimentally, and the droplet size distribution is predicted based on the maximum entropy principle. The results demonstrate that the spray cross-section in supersonic crossflow is approximately Ω-shaped, and the heights and widths increase linearly with an increase in the orifice diameter. The range of the cross-sectional droplet size is 0–30 μm, and the dominant influencing factor is the aerodynamic force. The cross-sectional droplet size depends on the transverse direction distance but not the orifice diameter. As the transverse distance increases, the peak intensity of the curves increases, and the droplet size becomes more uniform. The empirical relationship between the characteristic average diameter and the transverse distance at a specific distance (10–50 mm) from the orifice is obtained from the experimental data. The characteristic average diameter constraint method based on the maximum entropy principle provides an accurate prediction of the cross-sectional droplet size distribution. • Spray section of the liquid jet in the supersonic crossflow was approximately Ω-shaped. • The cross-sectional height and width increased proportionally with the orifice diameter. • More than 97% of the droplet size in the spray section was concentrated in the range of 0–30 μm. • A method can be used to predict the cross-sectional droplet size distribution. [ABSTRACT FROM AUTHOR]

Decarbonization in the aeronautical industry is a daunting challenge due to the lack of sustainable aviation fuel. Hydrogen offers a potential option for this problem, given that it is an energy-dense fuel capable of reducing climate impact. Along with it, the liquid hydrogen preheating and thermal protection bring inevitable safety issues. Currently, how to construct the optimal thermal management system (TMS) architecture and further clarify its principles is still unclear. Thus, this study compares different TMS architectures through optimized genetic algorithms (GA), and firstly points out the excellent architecture as well as its operating parameters. Strikingly, our results reveal the optimized algorithm can significantly improve efficiency and precision of the optimizing processes. Benefiting from adaptive population strategy and elite preservation strategy, optimized GA saves half the time compared to conventional GA, and the variance is 60 % smaller than that of conventional GA. Using optimized GA, the series-parallel scheme is determined as the best solution and it exhibits better advantages on lightweight. In addition, results also show the leading factor is air/helium heat exchanger area. In conclusion, we hope this work can provide guidance for the optimal GA, contribute to decarbonization, and provide more ideas for the optimization of hydrogen TMS in the future. • A systematic optimization model of TMS is developed by considering multi-factors. • Dynamic self-adaptation of population strategy is proposed in improved GA. • The optimal framework and controlling parameters of TMS is determined. • The mechanism of how the variety of factors synergetic evolve is found out. [ABSTRACT FROM AUTHOR]

In order to accurately control the gas temperature of a supersonic heat-airflow simulated test system, this paper introduces the working principle of the system, including the fuel supply subsystem and the air supply subsystem, establishes the mathematical model of the system, including the fuel supply subsystem, the air supply subsystem and the gas temperature, and analyses the characteristics of the system. Based on characteristics such as non-linearity, parameter time variations, disturbance and pure time delay exist in the system; a new cascade control scheme that can control both the fuel supply subsystem and the air supply subsystem is proposed to achieve an accurate control of the gas temperature. On this basis, the simulation and experimental studies are carried out by the proposed control algorithm. The results show that the proposed gas temperature cascade control algorithm can achieve a fast, stable and no overshoot control of the gas temperature in a wide operating range, and the control accuracy can reach at ±5°C.

The aerodynamic heating effect of the aircraft during supersonic flight has a great impact on the structural strength and thermal protection design. The aerodynamic thermal environment in the real state needs to consider the coupling of the external flow field and the structure and the influence of the boundary conditions of the inner wall surface. The S-A turbulence model is used to solve the Navier-Stokes equation. The coupling of the flow field and the solid wall is used to realize the coupled numerical analysis of the external flow field and the structural field. The coupling numerical study of the supersonic flow aerodynamic heating of three different airfoils is carried out, and the aerodynamic thermal effects of the inner wall surface of the airfoil under different thermal boundary conditions are compared. The results show that different airfoils have aerodynamic thermal effects similar to aerodynamic forces. The boundary condition of the inner wall considering convective heat transfer is the closest to reality; considering the fuel tank of the wing is full, the heat flux at the leading edge of the 3D wing can be up to 4 200 W/m2.

The problems of steady, inviscid, isentropic, irrotational supersonic plane flow passing a body with a small thickness ratio was solved by the linearized theory, which is a first approximation at and near the surface but fails at far fields from the body. Such a problem with far fields was solved by W.D. Hayes’ “pseudo-transonic” nonlinear theory in 1954. This far field small disturbance theory is reexamined in this study first by using asymptotic expansion theory. A systematic approach is adopted to obtain the nonlinear Burgers’ equation for supersonic far fields. We also use the similarity method to solve this boundary value problem (BVP) of the inviscid Burgers’ equation and obtain the nonlinear flow patterns, including the jump condition for the shock wave. Secondly, the transonic and hypersonic far field equations were obtained from the supersonic Burgers’ equation by stretching the coordinate in the y direction and considering an expansion of the freestream Mach number in terms of the transonic and hypersonic similarity parameters. The mathematical structures of the far fields of the supersonic, transonic, and hypersonic flows are unified to be the same. The similar far field flow patterns including the shock positions of a parabolic airfoil for the supersonic, transonic, and hypersonic flow regimes are exemplified and discussed.