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A modification in the ramjet engine, known as the Supersonic Ramjet Combustion, completes the combustion at supersonic speeds. Inside the combustion chamber, the mixing of fuels and the entire combustion procedure completes in milliseconds. So, the combustion chamber's flow physics is highly complicated. Through this study, the target is to improve the combustion efficiency inside the combustion chamber, thereby improving the performance of the scramjet engine. For this study, the performance of a strut-based flame stabilizer was analyzed in various positions, and the impact on energy efficiency and combustion performance is evaluated. SST k-ω model of turbulence is used to work out the Navier-stokes equation. To validate data from the experiment, a single strut configuration was considered for the next step, i.e., the comparison of the available data with a three-strut configuration. While considering the three-strut layout, in the x and y directions, secondary struts are placed, while the position of the primary strut remains constant. Only the flow from the primary strut is considered for the analysis of combustion performance out of three struts. From the obtained results, it is clear that the significant factor in improving energy efficiency is positioning the secondary strut in the present three-strut layout. The experiment results state that the combustion efficiency is boosted by 33.86% compared to a single strut combustor. It was also found that there was a considerable decrease in the level of unburned fuel, which makes the configuration more economical. The placement of struts plays a vital role, as any fault in the configuration makes it less efficient than a single strut layout. Other elements that boost combustion efficiency are expansion fans and shock reflection inside the primary zone of combustion and the secondary strut region. With the inclusion of secondary struts, the static wall pressure also escalated. A discrete flow was also observed on the combustion walls for some particular strut layouts. An increase in combustion efficiency was seen when both primary and secondary struts were placed near each other and also for reacting flow from the primary strut alone. There arises no need to supply fuel to secondary struts. So, overall fuel efficiency is improved, making the whole system energy-efficient. • Primary strut is considered for the analysis of combustion performance out of three struts. • The significant factor in improving energy efficiency is positioning the secondary strut in the present three-strut layout. • The combustion efficiency is boosted by 33.86% compared to a single strut combustor. • The intensity of combustion also increased which is evident from static temperature plots. • Shear shock layer enhances the mixing rate by intensifying turbulence phenomena. [ABSTRACT FROM AUTHOR]

In the paper (Theoret Comput Fluid Dyn 36:705–722, 2022), we analyzed acoustic receptivity of the boundary layer on a flat plate in Mach 6 flow at various angles of attack (AoA). It was shown that slow and fast acoustic waves passing through: a bow shock at AoA = - 5 ∘ , a weak shock induced by the viscous–inviscid interaction at AoA = 0 ∘ , or an expansion fan at AoA = 5 ∘ , excite dominant modes F and S in a small vicinity of the plate leading edge. The present paper extends this analysis to the cases of receptivity to entropy and vorticity waves. Similar to the case of acoustic receptivity, modes F and S of about equal amplitude are excited in a small vicinity of the plate leading edge. These modes propagate downstream in accord with the two-mode approximation model accounting for the mean-flow nonparallel effects and the intermodal exchange mechanism. Cross-comparisons of the initial amplitudes of excited modes help to evaluate the relative role of acoustic, entropy and vorticity waves in the second-mode dominated transition. [ABSTRACT FROM AUTHOR]

The results of calculations of the convective heating of a sharp plate placed in a supersonic flow at a speed of M = 6–8 are presented. The NERAT-2D code is used for calculations, which implements the Reynolds-averaged Navier−Stokes equations, together with Baldwin–Lomax and Prandtl algebraic models of turbulence. Good agreement is shown with the experimental data for convective heating in a turbulent boundary layer. The distribution of gas-dynamic functions is analyzed in different cross sections of the plate within the flow. [ABSTRACT FROM AUTHOR]

Direct numerical simulation and theoretical analysis of acoustic receptivity are performed for the boundary layer on a flat plate in Mach 6 flow at various angles of attack (AoA). Slow or fast acoustic wave passes through: a bow shock at AoA = - 5 ∘ , a weak shock induced by the viscous–inviscid interaction at AoA = 0 ∘ or an expansion fan emanating from the plate leading edge at AoA = 5 ∘ . The study is focused on cases where the integral amplification of unstable mode S (or Mack second mode) is sufficiently large (N ≈ 8.4) to be relevant to transition in low-disturbance environments. It is shown that excitation of dominant modes F and S occurs in a small vicinity of the plate leading edge. The initial disturbance propagates further downstream in accord with the two-mode approximation model accounting for the mean-flow nonparallel effects and the intermodal exchange mechanism. This computationally economical model can be useful for predictions of the second mode dominated transition onset using the physics-based amplitude method. [ABSTRACT FROM AUTHOR]

The present study investigates a novel control strategy to create a sweeping oscillatory jet using fluidic actuation principle. A supersonic underexpanded primary jet issued into a suddenly expanding duct is oscillated by giving cyclic pulsed secondary jet blowing in the top and bottom step regions of the suddenly expanding duct. The current study employs the bistable nature of jet expansion seen in a suddenly expanding duct for the jet oscillation. The cyclic perturbations (pulsed secondary jet blowing) given in the streamwise direction switch the direction of the asymmetric jet expansion of the underexpanded primary supersonic jet to the top and bottom wall alternatively, thereby creating an oscillatory supersonic jet. It is found that the switching of the direction of the primary jet to any one wall is caused by the Coanda effect produced by the secondary jet issued in the corresponding wall side. The results show that the primary jet quickly responds to the secondary jet perturbations and that the supersonic primary jet oscillation frequency nearly matches secondary jet perturbation frequency. A parametric study to investigate the jet oscillation characteristics with various secondary jet oscillation frequencies and mass flow rate ratios are also being carried out in the present study. It is found that the jet oscillation frequency increases with increase in the secondary jet pulsation frequency. The jet deflection is found to increase with increase in the secondary jet to primary jet mass flow ratio (blowing ratio) up to a critical limit, after which the increment in jet deflection ceases due to the restrictions imposed by the duct geometry. [ABSTRACT FROM AUTHOR]

U-shaped sonic booms associated with the dive maneuver of a supersonic vehicle are numerically simulated in this paper. Usual prediction tools for ordinary sonic booms are not suitable for such a waveform because of the breakdown of geometrical acoustics near the caustic. In this paper, nonlinear Tricomi equation approach is used near the caustic to take diffraction into account, whereas augmented Burgers equation analysis is applied away from the caustic where geometrical acoustics is applicable. The U-shaped waveforms predicted by this method are in good agreement with measured data. The effects of atmospheric turbulence are also considered by replacing one-dimensional analysis using augmented Burgers equation with multidimensional wave propagation analysis near the ground surface. Results with the atmospheric turbulence effects show better agreement with measured U-shaped waveform than ones without such effects. It is also found that the amplitude of the tail shock wave, which is much larger than the front shock wave, is more sensitive than the front shock to atmospheric turbulence. [ABSTRACT FROM AUTHOR]

In this study, the aerodynamic characteristics of the three-dimensional evacuated tube transportation (ETT) system based on the Reynolds-averaged Navier–Stokes κ−ω shear-stress transport turbulent model were investigated. The effects of two key parameters on the drag and flow topology of the ETT system, namely the travelling speed and ambient pressure in the tube, were studied. Compared with trains in the atmospheric environment without the tube (i.e., the open system), the ETT system shows considerable drag reduction with suitable operating parameters in the tube, particularly at a higher travelling speed range. The drag varying with the speed from subsonic to supersonic, shows various change trends at different speeds because of their distinct flow structures. The higher pressure in front of train head was observed to be reduced by choking, and a low pressure in the wake by expansion waves led to rapid increase in the drag and drag coefficient. The relationship between the drag and operating pressure was observed to be approximately linear for both the subsonic and supersonic speeds. [ABSTRACT FROM AUTHOR]

This paper investigates frictional dynamic contact mechanics of a functionally graded half-plane subjected to moving contact by a rigid flat punch possessing subsonic, transonic and supersonic speeds. The shear modulus along the half-plane is expressed by an exponential function, and the Poisson's ratio is assumed constant along the graded half-plane. Governing partial differential equations are derived based on the planar theory of elastodynamics. Boundary conditions are applied, and displacement fields in the graded half-plane are determined analytically. Formulation for the contact problem is reduced to a singular integral equation involving Cauchy singularity and a Fredholm kernel. Singular integral equation is solved numerically utilizing a suitable collocation technique. Contact stresses and normalized punch stress intensity factors are calculated for prescribed subsonic, transonic and supersonic speeds of the moving punch. It is expected that the results obtained by this study will help to understand the contact behavior and the surface failure mechanisms of functionally graded materials, especially at transonic and supersonic sliding speeds. [ABSTRACT FROM AUTHOR]

An experimental study is conducted on the interaction of a swept shock wave with a turbulent boundary layer. The shock wave is generated by a sharp unswept fin in a Mach 2 flow, where the strength of the interaction is varied from weak to moderate by changing the angle of attack α of the fin from 10 to 15 deg, which corresponds to a normal Mach number Mn of 1.3 and 1.4, respectively. Surface oil-flow visualization is used to study the topography of the interaction where surface flow features such as the upstream influence line and the separation lines are identified. By taking advantage of the quasi-conical symmetry of the flowfield, velocity field measurements are acquired in the conical reference frame for the interaction at the moderate interaction strength of Mn=1.4 at two locations from the fin apex. Flowfield features such as the λ-shock structure, slip line, and the separation bubble with the reverse flow are clearly visible in the velocity field. These results are also examined in the spherical coordinate frame, and good agreement in the spatial location of the critical features is found, providing direct quantitative, experimental evidence of quasi-conical symmetry of the flowfield above the surface. An examination of the velocity field downstream of the rear-foot of the λ-shock shows a region (a "streamtube," bounded on one side by the slip line emanating from the triple point) where the flow accelerates to transonic and supersonic speeds. This flow eventually turns toward and impinges upon the flat plate, a phenomenon referred to as an "impinging jet" in the literature. It is believed to be the principal cause of the high mean, unsteady pressures, very high heating, and skin friction coefficients that the present measurements confirm. An examination of the turbulence kinetic energy reveals a significant increase in amplitude in the separated flow region that is bounded by the flattened vortex under the λ-shock. [ABSTRACT FROM AUTHOR]

Acoustic analogy approaches can be used to express the sound radiation from a turbulent flow as the convolution product of a propagator and a source term. This paper presents an acoustic analogy-based analysis that properly accounts for circumferential variations of the propagator across the source correlation region in an axisymmetric jet. The analysis shows that including azimuthal source noncompactness effects leads to a formula for the acoustic spectrum in which each circumferential propagator mode couples only to its corresponding Fourier mode of the source term. This significantly affects the radiated sound field, because the lower-order propagator modes radiate much more efficiently than the higher-order modes, while the lower-order source modes usually contain significantly less energy than the higher-order modes. The resulting formula depends on the Reynolds stress autocovariance tensor, which must be accurately modeled in order to obtain realistic predictions of the sound field. A relatively simple, experimentally based model of this tensor is proposed and combined with Reynolds-averaged Navier-Stokes solutions to obtain predictions of the noise from a moderately supersonic cold round jet. The predictions are shown to be in good agreement with experimental data and the analysis is used to provide insights into the modal structures of the source and sound fields. [ABSTRACT FROM AUTHOR]

Peculiarities of the motion of one and two supercavitating strikers in water upon their simultaneous entry into water at speeds exceeding the speed of sound in the medium are considered. Conditions have been developed to ensure nondestructive entry of strikers into the water and stable movement along an aiming trajectory in the transonic speed range. A stable joint motion of two strikers in water at supersonic speed is obtained. [ABSTRACT FROM AUTHOR]

To investigate the complex multiphase flow of supersonic gas jets in still water, an underwater ignition experiment of solid rocket motor is conducted in the pressure tank, and the numerically simulated results of gas-water coupling for the same case are obtained by using Reynolds-averaged Navier-Stokes(RANS) method and volume of fluid(VOF) model. The results show that, after underwater gas jets achieve supersonic flow, the gas bubble appears to be cap-like due to the jet impact effect. Then the gas bubble gradually evolves into an ellipsoid-shaped gasbag with average axial expansion velocity of about 40 m/s. The flow structures inside the gas bubble are very complex, where two shear vortex rings and repeated shock cells exist. The interaction between the boundaries of gas jet and gas bubble may induce the instability of gas jet in the subsequent evolution stage. It is also demonstrated that there exists a part of high pressure in the flow field, and the pressure peaks gradually match with the environmental pressure in the oscillation process. The pressure of opening nozzle closure and the area of nozzle outlet are the two important factors that affects thrust peak value. [ABSTRACT FROM AUTHOR]

In this paper, 3-Dsupersonic flow around two types of wings is solved using a new algorithm for shock sensor calculation. A dual-time-stepping implicit method with 2nd-order accuracy is used for time integration of the equations. In each real time step, the non-linear system of equations is solved by iterating in pseudo-time, using a multi-step integration method. A cell-center finite volume scheme is applied to discretize the solution domain. Governing equations are discretized using 2nd-order central scheme of Jameson. Undesirable oscillations are prevented using artificial dissipation terms containing 2nd and 4th-order derivative terms. The second-order derivative term is proportional to shock sensor, which is a function of pressure gradient in general and is devised to capture shock waves correctly. Appropriate calculation of shock sensor is very important especially for the solution of 3-D supersonic flow on unstructured grids. In this study, a simple efficient algorithm is proposed for shock sensor calculation to stabilize solution in supersonic 3-D flows on unstructured grids. The new algorithm, implemented at an in-house code, is evaluated by comparison of its results with wind tunnel test data and upwindtype differencing scheme of Roe for a tailplane model tested at Royal Aircraft Establishment. The results show that supersonic flow with shock waves has been accurately captured. [ABSTRACT FROM AUTHOR]

Purpose – This study aims to decrease the aerodynamic drag of the body of revolution at supersonic speeds. Supersonic area rule is widely used in modern supersonic aircraft design. Further reduction of the aerodynamic drag is possible in the framework of Euler and Reynolds averaged Navier–Stokes (RANS) equations. Sears–Haack body of revolution shape variation, which decreased its aerodynamic drag in compressible inviscid and viscous gas flow at Mach number of 1.8 under constraint of the volume with lower bound equal to volume of initial body, was numerically investigated. Design/methodology/approach – Calculations were carried out in two-dimensional axisymmetric mode in the framework of Euler and RANS with SST model with compressibility correction equations at structured multiblock meshes. Variation of the radius as function of the longitudinal coordinate was given as a polynomial third-order spline through five uniformly distributed points. Varied parameters were increments of the radius of the body at points that defined spline. Drag coefficient was selected as an objective function. Parameter combinations corresponding to the objective function minimum under volume constraint were obtained by mixed-integer sequential quadratic programming at second-order polynomial response surface and IOSO algorithm. Findings – Improving variations make front part of the body become slightly blunted, transfer part of volume from front part of the body to back part and generate significant back face. In the framework of RANS, the best variation decreases aerodynamic drag by approximately 20 per cent in comparison with Sears–Haack body. Practical implications – The results can be applied for the aerodynamic design of the bullets and projectiles. The second important application is knowledge of the significance of the difference between linearized slender body theory optimization results and optimization results obtained by modern computational fluid dynamics (CFD) optimization techniques. Social implications – Knowledge about the magnitude of the difference between linearized slender body theory optimization results and optimization results obtained by modern CFD optimization techniques can stimulate further research in related areas. Originality/value – The optimization procedure and optimal shapes obtained in the present work are directly applicable to the design of small aerodynamic drag bodies. [ABSTRACT FROM AUTHOR]

Experimental research has been carried out of the processes occurring in a three-flow vortex tube operating in the humid air. Two types of swirling devices are investigated: single-pass spiral (snail) and multi-entry of five curved channels. An increase in the isentropic efficiency of a multi-entry swirling device by an average of 45% for a cold flow in the region μ = 0.10 ... 0.8 in comparison with a snail one has been experimentally confirmed. Three positions of the dividing gap were investigated at a distance of 4, 5, and 6 mm from the nozzle inlet. The drying efficiency of a snail-type swirling device is 10–15% better, and moisture removal through the separator is 10–25% better in comparison with a multi-entry swirling device. The vortex velocity distribution profiles are determined numerically using the SST turbulence model. It was revealed that an increase in the vortex velocity in the area of the swirling device by 15% made it possible to obtain an increase in the isentropic efficiency by 50%. A hypothesis has been put forward describing the mechanisms that realize the effect of breaking of liquid particles inside a vortex tube, which can reduce the efficiency of liquid separation by up to 25%. • A three-flow vortex tube is an effective device for the moisture separation. • Design parameters makes a significant influence on the vortex tube operation. • Multi-entry swirling ensures supersonic speed of the flow inside the vortex tube. • Multi-entry swirling provides 45% higher efficiency comparing to simple snail-type one. [ABSTRACT FROM AUTHOR]

A set of conceptual sketch for the near-space high supersonic (Mα = 3.5) unmanned aerial vehicle (UAV) is derived from the idea of integration of aerodynamic configuration and propulsion system. An accurate and straight-forward parametric geometry modeling for the near-space high supersonic UAV is essential to rapidly implement and evaluate the conceptual design. A geometry parameterization method for accurate description of the UAV configuration is proposed by use of shape function/class function transformation, Lagrange interpolation polynomial and B-spline curve. Based on the geometry parameterization, a code written in visual basic (VB) language is developed through the creation of macros from the CATIA application, which is used to implement the automatic generation of the 3-D geometry model of UAV. Three configurations of the high supersonic UAVs, including double swept wing, single swept wing and quasi-rhombus wing, are used to validate the method. The results show that accurate 3-D geometry models for all the UAV configurations can be rapidly and automatically created by the VB code. [ABSTRACT FROM AUTHOR]

The removal of nano-sized contaminant particles (CPs) from trenches formed on Si wafer is an important issue in the semiconductor industry. The particle beam technique, which is an advanced aerosol cleaning technique using nano-sized bullet particles (BPs) shot at supersonic velocity, has successfully cleaned CPs on flat surfaces down to 10 nm. In this paper, the same technique was applied to the removal of 10 nm AlO particles from micron-scale trenches using CO nano bullets. Trenches of 4, 10, and 20 μm depth were fabricated at 2 μm and 50 μm widths, with a maximum aspect ratio of 10. When nano-sized bullets were shot into the contaminated trenches, 10 nm contaminants deposited on the side walls, bottom surfaces, and corners were completely removed from a variety of trenches in just a few seconds. Cleaning efficiency greater than 95 % was attained. Size and velocity of the BPs were the important parameters affecting removal of CPs in the 10 nm size range from micron-scale trenches. [ABSTRACT FROM AUTHOR]

• The supersonic flow over reentry configurations with and without flap are analyzed. • The flap introduces streamwise and spanwise vortices and increases the base flow complexity. • At higher pitch angle, bow shock ahead of the nose and the oblique shock from flap interact. • On increasing the flap angle, a separation zone forms and grows in front of the flap. • At a constant Mach number, the axial force coefficient increases with altitude. Flap over blunt asymmetric reentry vehicles improves the aerodynamic stability and control. The present study has focused on the reentry aerodynamics of blunt-nosed conical body with flap configurations. Three dimensional, steady, viscous and compressible flow over the reentry body configurations were numerically analyzed by solving Reynolds averaged Navier-Stokes equations and SST K - ω turbulence model. The fundamental governing equations were discretized from the partial differential form to numerical analogue using the finite volume approach. Numerical simulations were carried out to investigate the flow characteristics of the three different reentry body configurations at different pitch angles, flap angles, the Mach numbers, and altitudes. For increment in the pitch angle from 0∘ to 6∘, the axial force coefficient is invariant, while the normal force coefficient linearly increases. It is found that the axial force coefficient is directly proportional to the flap angle and inversely proportional to the Mach number. Presence of flap introduces streamwise vortices and increases the flow complexity after the base to a large extent. [ABSTRACT FROM AUTHOR]