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Experiments were carried out to assess the effect of rear face geometrical modifications of a wall mounted rectangular cavity on pressure fluctuations at supersonic speed. Initially, tests were conducted on a base cavity of length to depth ratio of 3.6, in a 50 × 100 mm at a Mach number of 2, and a Reynolds number of 2 × 106 based on cavity length. Geometrical modifications to the rear face of the base cavity in the form of single ramp, double ramp and partial ramp circle were done to study their acoustic emission characteristics. Oil flow images revealed the complex nature of the flow inside the cavities. Time resolved flow visualization revealed attenuation of receptive interaction at the cavity leading edge, and minimized mass flux into the cavity by adopting these modifications. Unsteady measurements indicated reduction in the peak tones, broadband levels, and coherence. Among the modified geometries, double ramp configuration significantly reduced the overall sound pressure levels and the fluctuating pressures of the order of 10 dB and 50% respectively.

Exterior panels forming the skin of flight vehicles traveling through the atmosphere at high supersonic speed are often susceptible to the occurrence of limit-cycle type self excited vibrations called flutter. The development of methods for panel flutter control is desirable and it is the objective of this work by using piezoelectric materials. The nonlinear equations of motion for panel flutter using Von-Karman's large deflection plate theory and the quasi-steady aerodynamic theory are modified to consider the composite plate with the piezoelectric layer. The composite plate consists of two piezoelectric layers bonded perfectly to the upper and lower surface of the original panel layer. When a voltage is applied to the piezoelectric layers, an induced strain in the plate material takes place and consequently tensile in-plane load is obtained. These tensile loads shift the onset of flutter point to about twice the value of the critical dynamic pressure for the plate alone. The values of control voltage and the effect of piezoelectric material properties are investigated. The method gives good results and succeeds in suppressing the panel flutter at its operating condition.

Exterior panels forming the exterior skin of flight vehicles traveling through the atmosphere at high supersonic speed are often susceptible to the occurrence of limit cycle type self- excited vibrations called flutter. The development of methods for the prevention and control of flutter, without adversely affecting the weight and cost, is desirable and it is the objective of this work by using damping materials. The nonlinear equations of motion for panel flutter using Von-Karman's farge deflection plate theory are modified to include damping material layers. At first, a viscoelastic damping layer is considered to be perfectly bonded to the surface of the plate to investigate the extension damping treatment. Secondly, the viscoelastic damping layer over the plate is constrained by a third layer called constrained layer to investigate the shear damping treatment. The two types of damping treatment proved to be able to prevent panel flutter successfully at prescribed operating conditions. But for the same panel weight, the constrained damping layer technique is more efficient than the free damping layer technique.

Results of experimental studies of the round hydrogen microjet interaction with a coaxial air for supersonic speed jets efflux are presented in this work. It is shown that interaction of round and coaxial air jets at their supersonic efflux leads to appearance of a train small/large-scale and, on the contrary, large/small-scale supersonic cells on a resultant jet with growth and reduction of speed of its efflux, respectively. It is found that the result of the round and coaxial air jets interaction at their supersonic efflux does not depend on geometrical parameters of a round and coaxial exit nozzle.

Results of experimental studies of the round air microjet interaction with a coaxial hydrogen jet at its combustion for supersonic speed jets efflux are presented in this work. It is revealed that combustion of the coaxial hydrogen jet with growth of its speed efflux is accompanied by all scenarios, observed at study of the round and plane hydrogen microjets diffusion combustion. However, “bottleneck flame region” undergoes considerable geometrical deformations because of specifics of a flame of a coaxial jet. It is shown that “bottleneck flame region” is transformed from Y-shaped to spherical shape in the activity of growth of a coaxial jet speed efflux. It is found that a round air microjet interaction with a coaxial hydrogen jet at its combustion is accompanied by several new phenomena: existence of cone-shaped area a coaxial jet combustion near a nozzle exit; existence of small-scale supersonic cells on a resultant flame; absence of the hydrogen combustion efflux from combustion region of a coaxial jet near nozzle exit; flame-out from combustion region of a coaxial jet near nozzle exit that leads to hydrogen ignition downstream, its intensive combustion and sharp acoustic noise occurrence; existence of a turbulent flame, to its separation from a nozzle exit and transition to supersonic combustion of a resultant jet.

Experiments were carried out to bring out the effects of the single vertical tail (SVT) and the twin vertical tail (TVT) of a typical winged reentry vehicle on aerodynamic coefficients at supersoni...

High speed aerodynamic bodies are associated with high aerodynamic drag. Reduction of this drag is a necessary requirement, for which different active and passive control techniques have been proposed. In the present work, the concept of breathing blunt nose (BBN) for drag reduction at hypersonic speed has been applied to a long blunt nosed body at supersonic speed. Investigation on two breathing configurations has been done using CFD tools and effect of breathing on static aerodynamic coefficients particularly on axial force coefficient has been studied. The employed breathing configurations increase normal force coefficient and pitching moment coefficient by 5% and 10% respectively. The axial force coefficient is reduced by 3.3% and 4.4% for the two breathing configurations at α=0°. The Ca reduction increases linearly with angle of attack, and it reaches 6.25% and 6.4% at α=8°. The results of non-breathing body (NBB) have been validated quantitatively using experimental data; and the results of breathing configurations match qualitatively with those found in the literature.

A discharge device providing the ability to create an area in the interelectrode space with a controlled concentration of neutral particles due to the supersonic flow arrangement has been developed and implemented. The possibility to control the internal discharge parameters distribution by supersonic gas flow arrangement in a limited area of the glow discharge is shown experimentally.

The coupling of cavity noise and the structural vibration caused by the flow around the cavity at high speed will adversely affect the structure of the aircraft and threaten flight safety. Through the synchronous test method of noise and structural response, the test of acoustic/vibration coupling characteristics of the elastic cavity at supersonic condition in a wind tunnel was carried out, and the elastic cavity noise, structural vibration response characteristics and the coupling law were studied. The results show that the intracavity noise load mainly affects the low-order modal response of the cavity structure. Compared with the rigid cavity, the maximum position of the intracavity noise affected by the radiated sound wave is near the elastic boundary with strong vibration characteristics, and the noise peak frequency component is particularly affected.

Existing theories are reviewed and compared. From the Tangent Wedge formula a simple expression for the normal force on thin wings is derived and compared with some existing experimental data, which cover the whole supersonic-hypersonic range of Mach numbers.

The properties of a riblet surface at a Mach number of 1.25 are described. Direct measurement has been made of the surface shear stress on a riblet surface attached to a skin-friction balance mounted beneath the turbulent boundary layer growing along a wall of the RAE 2 ft×1 1/2 ft variable density transonic wind tunnel. Comparison of the data with that for a smooth wall indicates a reduction of about 7% in skin friction at optimum flow conditions. Misalignment of the riblet surface with the flow direction reduces this benefit and beyond 30° misalignment a drag penalty occurs. Analysis of the velocity profiles indicates that compressibility effects may be accounted for with the use of an appropriate transformation.

A closed form solution is given for the supersonic rolling of an elastic cylinder on a plane using the method of the preceding paper [1]. Tangential and normal forces are transmitted and Coulomb friction is assumed.

Experimental and numerical studies are carried out to validate the potential of opposing Plasma Synthetic Jet (PSJ) for drag reduction for a hemisphere. Firstly, flow field changes of opposing PSJ are analyzed by comparing the experimental schlieren images and simulation results in a supersonic free stream of Mach number 3. As PSJ is a kind of unsteady pulsed jet, the shock standoff distance increases initially and then decreases under the control of PSJ, which corresponds to the change of the strength of PSJ. Accordingly, the amount of drag reduction of the hemisphere increases initially and then decreases. It is found that there is a short period of “drag rise” during the formation of PSJ before the drag reduction, which is induced by the generation of normal shock waves and the area difference of the cavity wall of PSJ Actuator (PSJA). Secondly, the effects of five parameters, including exit diameter, discharge energy of PSJA, Mach number, static pressure of incoming flow and angle of attack, on drag reduction of opposing PSJ were studied in detail by using numerical method. It is found that the Maximum Pressure Ratio (MPR) has a significant impact on the average drag reduction for a configuration-determined PSJA. For the configuration selected in this study, the flow field of opposing PSJ shows typical Short Penetration Mode (SPM) in a control cycle of PSJ when the MPR is less than 0.89. However, the flow field shows typical Long Penetration Mode (LPM) at some time when the MPR is bigger than 0.89. Relatively better drag reduction is achieved in this case.

The present study proposes a segmented cooling-stream injection structure based on a certain coolant mass flow rate, and numerically investigates the effect of segmented cooling-stream injection on supersonic film cooling. The results indicate that without shock-wave impingement and with helium as the coolant, segmented cooling-stream injection can reduce the mixing between the mainstream and the cooling stream to produce better cooling performance than single injection, especially at larger coolant Mach numbers. However, with nitrogen as the coolant, the cooling effect of the segmented-injection system is very close to that of the single-injection system. Mixing at the impinging region is enhanced significantly when there is an incident shock wave. When the shock wave impinges between the two coolant inlets, segmented cooling-stream injection improves film cooling effectiveness in the midstream and downstream regions more than single injection because only part of the cooling stream undergoes the enhanced mixing effect of the shock wave. The advantage of segmented injection is reduced when the impinging region is behind the second coolant inlet. The further downstream the impinging region, the smaller the associated advantage.

The energy efficiency of a supersonic ramjet engine with detonation combustion of a hydrogen-air mixture in an axisymmetric Laval nozzle is considered. First of all, within the framework of the one-dimensional (1D) theory, it is shown that extreme combustion corresponding to the Chapman-Jouguet (C.-J.) deflagration, like constant pressure combustion, surpasses detonation in energy efficiency. However, this advantage becomes insignificant with an increasing the free-flow Mach number. Based on the results of the 1D theory, an axisymmetric Laval nozzle is constructed, which provides steady-state detonation combustion of a stoichiometric hydrogen-air mixture at an altitude of about 40 km and with freestream Mach numbers 9 and 12. The efficiency of detonation combustion in this case is about 35% instead of the maximum 65% according to the 1D theory.

The variable geometry supersonic inlet tends to decrease the throat area to reduce the Mach number upstream of the terminal shock, so as to reduce the flow loss. However, excessive Internal Contraction Ratio (ICR) exposes the inlet to a greater risk of unstart, which inevitably results in a process of increasing the throat area to aid the inlet restart. In the above throat regulation process, the inlet undergoes the start, unstart, and restart states in turn. In order to reveal the flow structure and mechanism of this process, a two-dimensional unsteady numerical simulation combined with a dynamic mesh technique were employed. The shock-on-lip Mach number of the studied inlet is 4.0 and the flight angle of attack is + 6°. Analysis was focused on the state with a freestream Mach number of 3.0. The results clearly show that the flow response hysteresis appears, and restart is only realized when the throat area is obviously increased as compared to that of unstart due to the historical unstart flow structure. In addition, three typical flow fields were analyzed, and it is found that the separation ahead of the inlet was the key factor affecting the hysteresis. Finally, unstart and restart boundaries of the inlet were discussed, and the factors influencing its deviation from the typical boundaries of dual-solution area were analyzed. The newly predicted unstart and restart boundaries are much closer to the CFD results.

Flow separation due to shock wave/boundary layer interaction is dominated in blade passage with supersonic relative incoming flow, which always accompanies aerodynamic performance penalties. A loss reduction method for smearing the passage shock foot via Shock Control Bump (SCB) located on transonic compressor rotor blade suction side is implemented to shrink the region of boundary layer separation. The curved windward section of SCB with constant adverse pressure gradient is constructed ahead of passage shock-impingement point at design rotor speed of Rotor 37 to get the improved model. Numerical investigations on both two models have been conducted employing Reynolds-Averaged Navier-Stokes (RANS) method to reveal flow physics of SCB. Comparisons and analyses on simulation results have also been carried out, showing that passage shock foot of baseline is replaced with a family of compression waves and a weaker shock foot for moderate adverse pressure gradient as well as suppression of boundary layer separations and secondary flow of low-momentum fluid within boundary layer. It is found that adiabatic efficiency and total pressure ratio of improved blade exceeds those of baseline at 95%-100% design rotor speed, and then slightly worsens with decrease of rotatory speed till both equal below 60% rated speed. The investigated conclusion implies a potential promise for future practical applications of SCB in both transonic and supersonic compressors.

Supersonic boundary-layer transitions triggered by a diamond-shaped roughness element were investigated by direct numerical simulation based on a finite volume formulation with a high-order minimum...

Suppression of shock-induced separation on all the walls of a rectangular duct using arrays of microvortex generators was experimentally investigated. The flowfield was analyzed using normal and in...

The discontinuity at the thermal protection system tiles covered on an aerospace vehicle is modeled as an open cavity to numerically investigate rarefied hypersonic flows over it under various flow control styles using the direct simulation Monte Carlo method. The influence of flow control on flow structures (e.g., shock wave and vortex) and aerothermal properties (e.g., temperature distribution and wall heat flux) is analyzed in detail. The flow structures and aerothermal properties of the cavity flowfield are both changed greatly by the square control block on the cavity opening line, such as a reduction of about 79.5% of the maximum of wall heat fluxes near the reattachment corner, while these changes are gradually weakened with the control block moving down. The control blocks with different shapes still have the ability to change flow structures and thus the heat flux coefficients over the cavity surfaces, but shape has little effect, and the square control block has the strongest wall heat flux compared with the circular, triangular, and wedged control blocks. For the supersonic jet with different directions, the leftward jet has the best performance in eliminating high temperature gas inside the cavity and over the downstream wall and reducing wall heat fluxes over cavity surfaces, followed by the downward jet. The main reason is that the leftward and downward jets can effectively prevent momentum and energy exchanges between freestream with cavity surfaces.

A study is undertaken to assess the dynamic control's effectiveness at two supersonic Mach numbers, 2.2 and 2.58 being considerably large Mach number being very significant in the design of High-speed aerospace vehicles. The relief available to the flow while the shear layer exits from the nozzle and enters the enlarged duct. The investigation area ratio is 4.84, and the NPR tested is from 3 to 11, covering the various level of expansion. The investigation was done for eight sizes of circular duct cases. Still, the results are discussed only when a significant change in the wall pressure is noticed during the experimentation. The dynamic control is located at a 6.5 mm radial distance from the primary jet's axis, with a 1 mm diameter generating at sonic Mach from the active flow control management. For lower duct length at NPRs, the wall pressure assumes large values. Later, with a rise in the duct section and NPR, a progressive decline in the pressure values is seen due to the decrease in adverse pressure levels and the intensity of the oblique shock waves. The duct flow mild oscillations are seen for duct section 3D, 5D, 6D, and 10D. The duct length portrays a crucial role while determining the stream development in the enlarged circular tube. When the duct's size is 6D and 10D, the low wall pressure is seen. For NPR = 5, the wall pressure takes moderate values due to the influence of the tube. At NPR = 7, 9, and 11, the pressure rise is limited within the reattachment zone, and ambient conditions are nearly attained within a short duct length. Despite nozzles, they are over-expanded flow control effects in a rise of pressure at Mach 2.58 for duct size 6D and 10D. For the duct size 10D, the attainment of ambient pressure is further advanced towards the upstream.

Selective metallization on glass surface technique has lot of potential applications in optoelectronics. Particularly, ultra-fine copper lines provide better optical transparency, which may enable to meet the demands of novel optoelectronic devices. In this paper, we propose a new manufacturing procedure to overcome the limitations of the traditional process and to achieve narrower copper lines on the glass surface. We, first, irradiate the glass surface by the femtosecond laser for more precise ablation, and then deposit the silver nitrate particles onto the irradiated region by using a supersonic particle deposition system. Copper lines with a linewidth of 1 µm were achieved by means of this new procedure, which is, to best our knowledge, the narrowest linewidth ever reported so far using laser direct writing and the subsequent electroless copper plating. We also demonstrate a highly transparent heater (i.e., 97% optical transparency) as an application of the proposed manufacturing approach.

With the increasing altitude and speed of the hypersonic vehicles, the temperature of the combustion chamber wall must be substantially decreased to ensure thermal and structural reliabilities. As a great supplement to the main regenerative cooling method without additional coolant, film cooling with onboard hydrocarbon fuel is widely used to decrease the wall heat flux. In this paper, a typical cylindrical film injection hole is employed to study the flow and heat transfer characteristics of gaseous film cooling with hydrocarbon fuel in the supersonic main flow, and the effects of transverse-non-uniform distributions of blowing ratio and injection angle are analyzed in detail. The results show that entrainment process between cooling films with different blowing ratios and injection angles would influence the film cooing efficiency to different degree, and proper combination of these two factors can improve the film cooling efficiency and decrease the flow rate of the cooling film. And the empirical correlation of supersonic hydrocarbon film cooling efficiency is established for the uniform blowing ratio cases. For the transverse non-uniform blowing ratio, the overall cooling efficiency decreases due to the spanwise and height-wise ejections between the two groups of cooling film. And for the cases with combination of transversely non-uniform blowing ratio and unequal injection angle, the film cooling effectiveness per blowing ratio would increase by 8.0%.

In this paper, the effects of non-isothermal oxidation on transient conjugate heat transfer of the cryo-supersonic air-quenching are investigated based on a double-layered oxidation kinetics model, while a unified conjugate heat transfer formula is developed to synthetically consider the near-wall turbulence, non-isothermal oxidation, and surface radiation. The comparison between numerical and experimental results are also presented to check the validity of the developed model. The results indicate that the film growth has some degree of inhibition to the conjugate heat transfer, in particular, the stagnation temperature increases linearly by about 5 K per 100 μm increase in film thickness, and the effective conjugate heat transfer coefficient in the stagnation region decreases linearly by about 55 Wm-2K-1 per 100 μm increase in film thickness. Moreover, the oxide film would have little impact on transient conjugate heat transfer when the near-wall velocity is higher due to the effect of viscous dissipation.

Nonintrusive laser diagnostics are used for the measurements of metastable nitrogen molecules in the lowest excited electronic state, N2(A3Σu+), in a nonequilibrium flow, blowdown supersonic wind t...

The mixing concept of fuel and air is the burning issue for hypersonic vehicles (scramjet) due to the less resident time of supersonic air in the combustion chamber. So far, significant research has been done for mixing enhancement and introduced different technologies; still, there is a lack of research for mixing improvement. Shock wave and shear mixing layer are the main parameters for investigating mixing criteria at supersonic speed. In this research, an innovative fuel injection strut has been designed to develop mixing enhancement by elevating multiple interactions between the shock wave and shear mixing layer. This new strut has been designed with the reference of the DLR scramjet combustor. From the reference of a wedge-shaped strut, a revolved (wedge shape – circular 3D) wedge strut has been modeled with the same fuel injection base points. This new strut's performance has been analyzed for mixing enhancement by visualizing the development of shock wave, shear-mixing layer, and their interactions. Three-dimensional numerical analysis has been carried out by solving the Reynolds-Averaged and Navier-Stokes (RANS) equations. A comparison of results has been made for the basic wedge and new strut and identified the increase in multiple interactions of the shock wave and shear layer, which leads to an increase in mixing enhancement. For the new strut, complete mixing has been achieved within a distance of 0.180 m with an average increase in mixing efficiency of 9% and increased pressure losses of 12%.

Modeling of velocities and temperatures processes distribution in the plasma-forming channel determining the design features and optimal parameters of the plasma torch nozzle is one of promising directions in development of plasma technologies. The aim of this work was to simulate the processes of velocities and temperature distribution in the plasma-forming channel and to determine the design features and optimal geometric parameters of the plasmatron nozzle which ensures the formation of necessary direction of plasma flows for generation of surface waves on the surface of a liquid metal droplet under the influence of the investigated instabilities.One of the main tasks is to consider the process of plasma jet formation and the flow of electric arc plasma. For obtaining small-sized particles one of the main parameters is the plasma flow velocity. It is necessary that the plasma outflow velocity be close to supersonic. An increase of the supersonic speed is possible due to design of the plasmatron nozzle especially the design feature and dimensions of the gas channel in which the plasma is formed. Also the modeling took into account dimensions of the plasma torch nozzle, i. e. the device should provide a supersonic plasma flow with the smallest possible geometric dimensions.As a result models of velocities and temperatures distribution in the plasma-forming channel at the minimum and maximum diameters of the channel were obtained. The design features and optimal geometric parameters of the plasmatron have been determined: the inlet diameter is 3 mm, the outlet diameter is 2 mm.The design of the executive equipment has been developed and designed which implements the investigated process of generating droplets of the micro- and nanoscale range. A plasmatron nozzle was manufactured which forms the necessary directions of plasma flows for the formation of surface waves on the metal droplet surface under the influence of instabilities. An algorithm has been developed for controlling of executive equipment that implements the process of generating drops of micro- and nanoscale range.

The aerodynamics of re-entry vehicles has always been in the forefront among space research. With the increasing interests in space exploration, it is important to understand the aerodynamics of spacecrafts and identify an optimum design to overcome the forces due to the earth's atmosphere. In this paper, Computational Fluid Dynamic (CFD) analysis has been carried out on ballistic bodies with a single, double and triple aero-disk spikes. The study also involves the presence and absence of a blunt tip to aid the analysis of flow behaviour on these bodies under two scenarios namely zero-degree angle of attack and at supersonic speed (Mach 2.0). The aerodynamic drag and surface heat transfer effects of these spikes on three different blunt body configurations has been studied. The shape of the blunt bodies vary based on the nose bluntness factor. The flow patterns of these configurations were three-dimensionally visualized via Schlieren-like, Mach and Pressure contours. It was observed that the usage of aero-disk spike reduced the aerodynamic drag by 45%–60% when compared to the bodies without spikes. The configuration labelled 3221 showed the maximum effect with an average drag reduction of more than 55%. There was a remarkable drop in the stagnation point value of surface heat flux with the addition of all five spike system. An efficient performance of localized peak heat fluxes was also noticed upon the addition of double and triple spike systems. In conclusion, it has been observed that the usage of double and triple spikes on the 3221 blunt body has significantly improved its aerodynamic drag and its surface heat flux.

Micro-ramps for shock train control were investigated experimentally in a Mach 1.85 supersonic isolator flow using three device heights. A mechanical flap was mounted downstream and linearly choked...

An experimental study concerning an underexpanded, screeching, Mach 1.5 jet operating at a stagnation-to-ambient pressure ratio of 4.4 is presented. Experimental data were acquired from high-speed ...

A more detailed understanding of the role of N2 in very low Earth orbit (VLEO) on atomic oxygen (AO)-induced material erosion is necessary for future development of the sub-LEO region. Accordingly, a laser-detonation source capable of capturing two different beams is developed. A dual-pulsed supersonic valve (dual-PSV) system is designed to form two beams in one nozzle configuration. It is demonstrated that the dual-PSV is capable of forming two atomic beams at individual energies with various composition ratios. This system is successfully applied for a ground-based simulation of the world’s first material erosion experiment in a sub-LEO aboard Super Low Altitude Test Satellite (SLATS).

The supersonic mode in hypersonic flows is studied numerically. The radiation characteristics of this mode in a Ma=6 flat-plate boundary layer with cooled walls are focused on. The linear stability...

Boundary layer development in a supersonic inlet is studied, focusing on the flow over the compression ramp and the ramp–isolator junction. Two designs of the junction, a backward-facing step (BFS)...

During atmospheric reentry at hypersonic speed and during very-high-speed flight in the atmosphere, vehicles experience extremely high thermal loads for relatively long time. To handle this problem...

This paper reports an experimental study on the supersonic jet surface flow structure visualization and shear stress field measurement issuing from a rectangular nozzle with extended shelf. The evolution of the near-field surface flow structures with an increased Nozzle Pressure Ratio (NPR) is successfully captured by the surface oil flow, infrared detection technology, and the Shear-Sensitive Liquid Crystal Coating (SSLCC) technique. Results reveal that under smaller NPR, the wall flow structure is similar to that of a jet without the extended shelf i.e., clean jets, and this is caused by insufficient effect on the boundary layer. However, at higher amplitudes of NPR, there exists a significant effect of the boundary layer, as a near triangular separation forms on the trailing edge of the Mach stem due to the adverse pressure gradient, which is visualized for the very first time in this paper. Furthermore, the vector field of shear stress is measured quantitatively by SSLCC technique. Results shows that the magnitude of shear stress heightened with NPR increasing, and the directions of shear stress changes across the shock wave and expansion fans. In addition, surface streamlines measured by SSLCC is significantly consistent with the streamlines visualized using the oil flow technique.

In this study, the flutter analysis of a rotating pre-twisted functionally graded graphene nanoplatelets reinforced composite (FG GPLRC) blade under supersonic airflow is investigated. The pre-twisted blade is multi-layered and reinforced with graphene nanoplatelets (GPLs) evenly distributed in each layer while the GPL weight fraction changes from layer to layer through the thickness direction. The effective Young's modulus is determined by Halpin-Tsai micromechanical model while the Poisson's ratio and mass density are predicted by Voigt's rule for GPLRC layers. According to assumptions of the first-order shear deformation theory (FSDT), the first-order piston theory and shell theory, the dynamic model of rotating pre-twisted GPLRC blades subjected to supersonic flow is developed. Meshless the improved moving least-square Ritz method (IMLS-Ritz) is used to derive the discrete dynamic equations of rotating pre-twisted GPLRC blades under aerodynamic load. The accuracy of the IMLS-Ritz method for this problem is validated by comprehensive convergence studies and careful comparison studies. A detailed parameter investigation of the effects of GPL distribution configuration, rotation velocity and geometrical parameters on flutter behavior characteristics of GPLRC blades is systematically conducted.

In a Supersonic combustion ramjet (Scramjet) engine, combustion occurs at supersonic velocity as incoming air remains supersonic. Scramjet engines have complex flow phenomena taking place inside the combustor. In a scramjet combustor, mixing and combustion should take place within few milliseconds. In the present study, two additional trailing struts with no fuel injection are placed a short distance downstream of the fuel injection strut. Effect on combustion performance using these strut-based flame stabilizer configurations are assessed. Reynolds averaged Navier-Stokes equations are solved with turbulence model and species transport equations. Validated results of the single strut are compared with the different strut configurations. It is found that the placement of trailing struts have a vital role. High-pressure zones, shock reflections, recirculation region, and expansion fan in between combustion region and the trailing struts enhance the combustion efficiency. Strut configuration introducing high-pressure regions outperformed other strut configurations due to the formation of a strong recirculation region in the core combustion region. Combustion efficiency is found to have a maximum improvement of 67.14% compared with single strut-based flame stabilizer. Total pressure loss also increased due to the introduction of additional struts.

The major hurdle in augmenting the performance of supersonic engines is the appropriate distribution of injected fuel within the supersonic chamber. That's why SST k − ω model is used to explore the mixing characteristics of a transverse jet injection into a supersonic airstream. In this aspect, initially, the performance of the supersonic combustor is analysed using hydrogen and ethylene fuel in terms of mixing efficiency. Our results suggest that the flame holding capacity of hydrogen-fuelled combustor is better as than an ethylene-fuelled combustor. Additionally, to strengthen the mixing of fuel with supersonic air-flow, a novel shock generator is introduced and placed at different locations of the combustor. The flow-structure is studied comprehensively using two single shock generator and two dual shock generator to conclude the best possible geometry for improved performance. The results reveal that the fuel penetration depth of geometry 3 with dual shock generator has increased considerably. Moreover, geometry 3 offers enhanced mixing between the supersonic air and fuel compared to remaining models with the formation of a strong vortex pair and recirculation zone near the injection slot.

The detailed mechanism of supersonic mixing enhancement by a novel device is investigated computationally based on the Navier-Stokes equations. The device is a wall-mounted three-dimensional cavity, several parts of the rear face of which are cut out, and the lower wall of the cut-out part has a slope directed toward the injector from which the injectant is discharged. The flow structures and behaviors observed in the computational results are confirmed in the experimental results. Experimental information about the high-frequency (several tens of kHz) pressure oscillation in the cut-out part is acquired in this study using a pressure-sensitive paint (PSP) responding to rapid change in pressure. The computational results reveal that a hairpin-like vortex is formed inside the cavity. The vortex works on the shear layer across the cavity to deflect it upward. The deflected shear layer shields the injectant from the primary flow, which causes the injectant to penetrate deep into the primary flow. The computational results also clarify that the expansion and compression waves produced in the rectangular part of the cavity are delivered effectively to the injector along the slope of the cut-out part, which is confirmed in the PSP measurement results. The delivered waves actuate the jet issuing from the injector to fluctuate up and down periodically with a large amplitude.

We investigated the oxidation of oxygen vacancies at the surface of anatase TiO2(001) using a supersonic seeded molecular beam (SSMB) of oxygen. The oxygen vacancies at the top surface and subsurface could be eliminated by the supply of oxygen using an SSMB. Oxygen vacancies are present on the surface of anatase TiO2(001) when it is untreated before transfer to a vacuum chamber. These vacancies, which are stable in the as-grown condition, could also be effectively eliminated by using the oxygen SSMB.

Efficient mixing is of great importance for engineering applications of hypersonic propulsion systems. With supersonic air flowing through the interior of combustor in milliseconds, rapid mixing of airstream and fuel becomes one of the great challenges, and it is imperative to employ mixing enhancement strategies to improve the system performance. In present work the splitter plate jet assisted mixing (SPJAM) strategy is proposed and direct numerical simulations are conducted to reveal enhanced-mixing mechanisms. Both instantaneous fine structures and turbulent intensity distributions are obtained to analyze the flow dynamical behaviors. Mixing layer thickness growth rate and structure topology properties affected by SPJAM are researched as well. The results indicate that the introduction of SPJAM can significantly promote mixing by means of the newly found T-shaped structures and its dynamics. The interaction of T-shaped structure and inherent Kelvin-Helmholtz vortex leads to the breakdown of flow structures, which can obviously increase the interface of upper and lower streams. Turbulent intensity analysis suggests that with the effect of SPJAM, the region where turbulent activity exists is notably extended, and the peak values of the fluctuations are much larger. The evolution of transversely-integrated turbulent kinetic energy (IK) indicates that rapid increase of IK appear in the near field of the flow. Spatial correlation analysis results indicate that vortex size is notably increased with the introduction of SPJAM, especially for that with the strategy set at the front of the splitter plate. Through exploration of the enhanced-mixing mechanisms, present preliminary work indicates that the proposed strategy is a good candidate for efficient mixing, and in the future more detailed work including three-dimensional simulations concerning the strategy optimization should be conducted.

A comprehensive study was performed on the atomization of molten stainless steel in a high-pressure gas atomizer (HPGA). Computer simulations were carried out using ANSYS-FLUENT to model the supersonic, turbulent flow of the atomization gas. Particle tracking and solidification were also included in the model to determine the fate of molten particles within the atomizer. Experiments were performed to investigate the effect of varying gas pressure and nozzle diameter on particle size distribution, melt flow rate, and pressure gradients on the surface of the nozzle. It is shown that increasing the atomization pressure leads to a finer powder size distribution, but a lower powder throughput. The results and conclusions provided in this work provide a valuable insight into how various operating parameters can affect the performance of HPGAs for metal powder synthesis. The findings can be used to improve the design of these systems in terms of efficiency, throughput, and product quality.