The interaction of external shock waves with cavity shear layers plays an important role in determining the aerodynamic behaviour of high-speed cavity flows. In this study, a numerical investigation is carried out to examine shock impingement on a supersonic open cavity, with focus on the effects of the vertical position of the shock generator and the freestream Mach number. An off-body shock generator is placed at different heights above the cavity opening to produce oblique shocks that interact with the cavity shear layer. Two-dimensional steady Reynolds-Averaged Navier-Stokes simulations are performed using a density-based solver with an appropriate turbulence model. The flow fields are analysed using Mach contours, pressure contours, shock structures, and surface pressure distributions. The shock generator height is varied at three locations relative to the cavity depth, and simulations are carried out over a range of freestream Mach numbers. The results show that lower shock-generator positions cause stronger interaction with the cavity shear layer, leading to greater shear-layer deformation, stronger internal shock reflections, and higher surface pressures. As the generator height increases, the interaction weakens and the flow gradually approaches the baseline case without shock impingement. Increasing the freestream Mach number amplifies these trends, especially for lower generator positions.

The present research includes a comprehensive computational fluid dynamics (CFD) investigation to enhance the scramjet combustor performance. Although scramjet propulsion systems have a lot of potential for use in hypersonic flight, they face substantial technological difficulties when operating effectively at supersonic speeds. The present study’s main objective is to analyse a dual-strut and cavity-based model of a scramjet combustor in order to boost its effectiveness. To analyse the intricate flow and combustion processes that occur inside the scramjet combustor, the study employs a high-fidelity CFD technique. To precisely depict the complicated features of supersonic combustion, a two-dimensional, compressible Reynolds averaged Navier-Stokes solver with turbulence modelling is used. By including the struts and cavities in the design, various modifications are made to the standard DLR model, and the flow is then analysed. In the current simulations, the air velocity at the strut and species fraction at the outlet of the combustor are primarily observed, which aids in understanding of the mixing time and combustion efficiency of the combustor. The dual strut with cavity creates the recirculation region that could enhance the fuel air mixing of the combustor. The distribution of hydrogen in the lateral direction of the flow is improved by the shock wave interaction from the cavity compared to the baseline strut design. The findings of the present study advance our understanding of scramjet combustor performance and offer insightful information for upcoming design advancements.

Scramjets are considered one of the important propulsion systems for realizing hypersonic flights and have attracted a lot of research interest over the past few decades. The present study focuses on conducting a comprehensive numerical investigation on a scramjet inlet-isolator at Mach 5 with the motivation to improve efficiency. The study investigates a two-dimensional scramjet inlet-isolator, analyzing both external and internal flow characteristics by using the commercially available tool Ansys-Fluent. Numerical simulation is the first step to incorporate the passive bleed mechanism into the inlet design features with alterations aimed at improving airflow regulation. The subsequent investigation is the analysis of the effect of a number of cowl deflection angles on performance of the positions. An extensive evaluation is presented on the effect of varying cowl deflection angles on the inlet isolator efficiency. A combined approach is also adopted in the study whereby the most promising configurations from the bleed mechanism and cowl deflection angle studies are incorporated into a single geometry. This aims at increasing the inlet isolator efficiency and the pressure recovery. The aim of this integrated strategy is to develop an inlet geometry that is optimal for performance of the engine and pressure recovery capability enhancement. The study analyzes ten different cases, and it is noted that out of these, the 4-degree cowl deflection angle is the most efficient in enhancing the total pressure efficiency.

To minimize forebody drag in high-speed flying vehicles such as missiles and rockets, contemporary research has focused on computational methods to analyze drag reduction strategies. This study investigates the efficacy of an intermediate aerodisk mounted on a sharp-tip spike at a Mach number of 2.0. Through a parametric analysis, variations in aerodisk size and location on the spike stem are explored. Results indicate that reducing the size of the intermediate aerodisk to 3 mm maintains identical reattachment shock strength but leads to higher pressure values at the transition from separation shock to reattachment shock. The model with an expanded 5 mm aerodisk size exhibits the second-lowest peak pressure coefficient for reattachment shock, suggesting improved flow recirculation and lower heating levels. Conversely, a 6 mm aerodisk size increases reattachment shock pressure but enhances flow recirculation, impacting total drag. Overall, the study concludes that an intermediate aerodisk, particularly with a 5 mm diameter, provides an optimal configuration for drag reduction before flow separation.

Comparative Computational Analysis of NATO 5.56 mm, APM2 7.62 mm and AK-47 7.82 mm Bullet Moving at Mach 2.0 in Close Vicinity to the Wall Various rifles require unique bullets. Each bullet has its capability, speed, and impact on the target. In metropolitan warfare, several bullets are shot close to the solid walls. These near walls affect the pressure distribution over the entire asymmetric bullet. The influence of a reflected shock depends on the angle at which it was reflected and the altitude from the ground to the body of the bullet. The current research emphasizes three bullets of varying diameters used in different types of guns. The first bullet is of NATO 5.56 mm, the second is APM2's 7.62 mm bullet, and the third is a 7.82 mm bullet from an AK-47 rifle. For 2-D steady computations, the supersonic speed of Mach 2 is considered to analyze the flowfield across all three bullets. The heights of the bullet are taken considering the height-to-diameter ratios (h/D ratio) from 0.5 to 3.0. The Mach contour drawn from the numerical simulations is used to analyze the flowfield, and aerodynamic coefficients like lift, drag, and moment are also plotted to analyze the ground effects on the projectile. The comparative analysis showed that the trend of shock wave reflections was similar in the bullets till h/D of 1.5. The APM2 bullet experienced maximum drag, followed by AK-47's 7.82 mm and NATO's 5.56 mm bullet. The 7.82 mm bullet experienced maximum lifting force at h/D = 1.0 due to its larger surface area than the other two ammo. The 7.82 mm bullet experienced a nose-up moment, whereas the other two faced a nose-down moment. As the altitude of the bullets from the ground increased, the ground effect appearing on the bullets reduced. The present comparative analysis research shows that it is suitable to fire an AK-47 bullet from h/D greater than 2.0 and the other two bullets from an altitude greater than or equal to h/D of 3.0.

Humankind consumes different forms of energy for life and growth. Electric energy is one of them. Demand for electric energy keeps growing as the application area of electric energy keeps increasing day by day. To meet this huge demand and to save natural recourses, there is research going on to generate electrical energy from renewable energy resources like solar energy, wind energy, etc. Energy harvesting technology is an emerging area that gives the opportunity to increase the efficiency of reusable energy generation by using recent technology like artificial intelligence, machine learning, the Internet of Things, etc. The dissipated energies like electromagnetic waves, heat energy and vibrations are converted to electric energy. Artificial intelligence can be deployed along with different sensors like piezoelectric sensors for energy harvesting. This harvested energy can effectively be useful in different sectors like automobiles, domestic and industrial applications, etc. This book chapter summarizes the characteristics of energy harvesting with the help of AI/ML, recent developments in this area and major challenges, along with the future scope of development. Application of existing artificial intelligence technology in the field of energy harvesting is explained and the possible use of existing state-of-the-art AI/ML technologies for prediction 2and increase the efficiency of energy harvesting technology is also suggested in the present book chapter.

Supersonic intake experiences various types of shock-shock and shock boundary layer interaction patterns along the inlet and isolator sections. Also, the flow path consists of small, protruded objects which cannot be avoided from the design point of view. Such protrusions induce not only the aerodynamic and thermal load on the surface of the aerodynamic bodies. In the present work, computational analysis has been performed at Mach number 2.0 to investigate the shock wave reflection pattern generated by a set of triangular-type protrusions placed along the scramjet isolator and its effect on the performance parameters. The isolator is a constant area duct between the inlet section and the combustion chamber of the scramjet intake. For the computational analysis, a duct was considered with constant length and height in which two wedges were kept at in the middle of the duct facing each other, then the flow was passed over it. The oblique shocks generated due to the presence of the wedges, the shock type has been determined past the wedges whether the deflection contributes to Mach reflection or Regular reflection. Using theoretical analysis, the pressure deflection diagram has been obtained by varying turn angles with a constant freestream Mach number and correlated with the data obtained from numerical simulations. The comparison of shock polar suggests a small transition region between Mach reflection and regular reflection phenomenon. It was also observed that the flow gets detached before the maximum deflection angle due to the incoming boundary layer.

In the present study, shock waves produced over different blunt bodies moving at Mach 2.0 at zero angle of attack have been captured and analysed using computational investigations. Forebodies of distinct bluntness (hemisphere, ellipse, and ogive) are adopted, and the effect of the forebody geometry on the flow parameters is investigated. The flow parameters such as shock stand-off distance and shock strength have been captured over various forebody geometries and correlated with the forebody drag coefficient. With the reduction in the forebody bluntness, the shock stand-off distance and the shock strength are found to be decreasing, resulting in a decrease in the forebody drag coefficient values. The effect of the free-stream supersonic Mach number over a typical blunt body (hemisphere) was also investigated. With increasing freestream Mach number, the shock wave generated ahead the blunt hemisphere moved closer to the forebody and an increase in the forebody drag was observed.

Frontal cavity consists of complex unsteadiness which causes bow shock instabilities when the body is moving at supersonic or hypersonic speed through a fluid. These complex unsteady flow fields are an important practical concern in the aerospace applications. Understanding this flow field around the frontal cavity will help us to know more in depth about these instabilities to create effective control methods to avoid structural damages. The main objective of the present research is to analyze the bow shock instabilities at different Mach number around the frontal cavity at supersonic and hypersonic speeds by using the numerical concept of Detached Eddy Simulations (DES) in two dimensional axisymmetric domain. Three different freestream Mach numbers (2, 4 and 6) were considered in the present investigation in order to study the effect of Mach number on flow unsteadiness. It was observed that an increase in Mach number, the frequency of most dominant mode (bow shock pulsation) increases significantly. In addition to that, additional low amplitude sub-dominant modes were observed for all the cases of freestream Mach numbers studied in the present investigation.

The bullet is shot near the wall or the ground during urban warfare. This nearness leads to a pressure distribution on the bullet's entire body, which is asymmetric. For every case of difference in the height of the ground/wall proximity, the shock reflection angle changes, and a particular shock wave's effect on the bullet varies. In the present study, the flowfield around a 7.82-mm bullet of an AK-47 Rifle moving at a supersonic speed of Mach 2 near the ground is studied using Computational Fluid Dynamics. Computational simulations have been carried out for the bullet moving at different heights from the nearby wall. The level of impact on the projectile is illustrated from the region of nearest influence to a ground distance which is five times the bullet's diameter. Ground effects from height to diameter ratios (h/D) of 0.25 to 5 are showcased, describing the influence on overall flowfield, pressure coefficient distributions, lift drag, and moment coefficients. The wake region flow field is also analyzed. It has been observed that a detached bow shock wave is generated at the tip of the bullet, which increases the drag coefficient experienced by the bullet. Generation of lift is the maximum for certain h/D cases like 1.0, 1.5, and 2.0 as the shock reflections from the ground hit the bullet. Further increasing the altitude from the near-wall reduces the lift and drag forces acting on the bullet due to the lesser ground effect. From the present investigation, the heights above the h/D ratio of 2.0 are suitable for firing the shot.

This study is focused to develop the understanding of bow-shock instability observed around a frontal hemispherical shell in supersonic flow through numerical simulations. The frontal bow shock instability is observed in many situations such as canopy of supersonic decelerators, wind tunnel testing of rigid concave hemispherical shell in supersonic and hypersonic flows. In previous numerical studies, the small amplitude fluctuations of bow shock in front of hemispherical shell are well captured, but rapid large amplitude fluctuation in bow shock has caused failure in simulation. Hence, the principal objective and novelty of the study is to capture and analyse the experimentally observed 'large amplitude' shock unsteadiness in front of hemispherical shell in a supersonic flow of Mach 4 using the recently emerging numerical concept of Detached Eddy Simulations (DES).

In aviation, open cavities have a critical role in influencing performance characteristics. Components such as combustion chamber, landing gear, and weapon bays contain a cavity influenced by the shear layer. Supersonic conditions of cavity flows are a prerequisite for the configuration of missiles and future aerospace vehicles. In the present research, the stream passing above a rectangular open pit containing a shear layer is computationally studied after the impingement of an oblique shock wave generated by a wedge type shock generator. The cavity duct investigated in the present study has a length to depth ratio of 3 at zero degrees angle of attack. Investigation on the influence of flow characteristics such as the magnitude of drag, pressure distribution, and recirculation within the duct has been performed and analyzed for cases in the presence and absence of shock generator at Mach 3. No particular disruption of the shear layer occurs in the absence of an oblique shock. However, a robust recirculation in the shock impingement cases is observed. The shear layer lifts up due to the thrust force generated in the open cavity on the impingement of a shock wave. A strong rear shock is generated at the rear wall separation point of the open duct. Further, the effect of the rise in the strength of the impinging shock on the entire flow over the cavity is also investigated. As the strength of impinging shock increases, it was observed that the magnitude of pressure distribution on the cavity surface also increases.

Shock-related unsteadiness over axisymmetric spiked body configurations is experimentally investigated at a freestream supersonic Mach number of 2.0 at 0 $$^\circ $$ angle of attack. Three different forebody configurations mounted with a sharp spike-tip ranging from blunt to streamlined (flat-face, hemispherical, and elliptical) are considered. Steady and unsteady pressure measurements, short-exposure high-speed shadowgraphy, shock footprint analysis from $$x-t$$ plots, and identification of dominant spatiotemporal modes through modal analysis are carried out to explain the unsteady flow physics. The present investigation tools are validated against the well-known events of ‘pulsation’ corresponding to the flat-face case. The hemispherical case is characterized by the formation of a separated free shear layer and associated localized shock oscillations. The cycle of charging and ejection of fluid mass from the recirculation zone is identified to drive the flow unsteadiness. Such an event triggers the movement of the separated and reattachment shocks in the opposite direction with respect to each other (out-of-phase shocks motion). In the elliptical case, the overall flow field resembles that of the hemispherical case, except with dampened unsteadiness. The value of the cone angle ( $$\lambda $$ ) associated with the recirculation region is found responsible for the fluctuations caused by the charging and ejection of fluid mass. Thereby, it controls the extent of out-of-phase shocks motion. In the elliptical case, shock unsteadiness is reduced as $$\lambda $$ is observed to be smaller. Based on the gathered results and understanding, the reduction in unsteadiness associated with the aerodisk mounted on the hemispherical forebody is demonstrated and explained via the almost complete elimination of the out-of-phase shocks motion.

A computational study is carried out by adopting Two-Dimensional Double Bump at supersonic air intake for Mach number 2.2. Numerical simulation has been made with an Explicit coupled solver with K-ω turbulence model. The improvement in flow and overall performance of supersonic intake is observed with adoption of Double Bump. It has been observed that the pressure recovery is increased by installing Double Bump. Computational study revealed that the low-energy air region near the wall boundary layer has minimized for the case of double bump compared with that of the single bump air-intake.

Flow over a conventional delta wing has been studied experimentally at a subsonic flow of 20 m/sec and the flow field developed at higher angle of attack varying from 10° to 20° has been captured. A vortex generator is mounted on the leeward surface of the delta wing and its effect on the flow field is studied. The set of wing tip vortices generated over the delta wing is captured by the oil flow visualization and the streamline over the delta wing surface captured with and without a vortex generator are compared. Based on the qualitative results, the effect of the vortex generator on the lift coefficient is anticipated. Further, force measurement is carried out to quantitatively analyze the effect of vortex generator on the lift and drag coefficient experienced by the delta wing and justify the anticipation made out of the qualitative oil flow visualization tests. In the present study, the effect of mounting of a vortex generator is found to be minimal on the lift coefficient experienced by the delta wing. However, a significant reduction in the drag coefficient with increase in angle of attack was observed by mounting a typical vortex generator.

A Computational Fluid Dynamics (CFD) analysis was conducted on an axisymmetric boat-tailed afterbody operating at an angle of incidence of zero degrees. The boat-tail drag can significantly affect the performance of overall propulsion system in rockets and missiles. These computations show a very complex regime when the base region flow connects with the flow from the nozzle and the ambient air, with strong adverse pressure gradients and shock induced separations in the flow field. Computations have been carried out at transonic Mach number 0.9 with nozzle pressure ratio (NPR) of 4. The Shear Stress Transport (SST) turbulence model has been used in calculations to study the flow difference in pressure and velocity contours as it provides more accurate results with the data obtained from experiments. Nevertheless, the boat tail surface pressure coefficient for a straight cylinder and a tapered cylinder (or boat tail) has been plotted and analyzed. The boat tail configuration model demonstrated a better performance in terms of drag experienced by the system.

Computational Fluid Dynamics was used to study the flow around an AK-47 Assault Rifle© 7.82-mm bullet in proximity to a near wall at Mach 2. The computational methodology was validated against the previously existing computational data. In metropolitan battles situations, the projectiles are fired in close proximity to the ground. The presence of ground prompts an asymmetric pressure distribution over the surface of the bullet. The reflection of shock waves after hitting the wall is considered particularly. From the earliest stage to the influence around the shot; the level of impact is subjected to the ground clearance. Streamlines across the projectile and related surface pressure distribution acquired from computational iterations are described. Clearances with 0.5 and 1 height/diameter ratio were studied to examine the transitions in aerodynamic coefficients like lift, drag, and moment.

A parametric analysis was conducted to determine the flow over a missile released from the weapon bay cavity passing at a supersonic speed of Mach 2 in comparison with Mach 5. A weapon bay cavity with specified dimensions and a missile at 2.75m distance from bottom cavity wall was tested in compressible flow RANS solver. Detailed information of complicated aerodynamic characteristics like pressure variation, density variation and Mach number was obtained through CFD simulations to study the variation in flow physics over missile. Pressure fluctuation experienced by missile surface is plotted as coefficient of pressure (Cp) data over a missile kept in supersonic flow of Mach 2 and Mach 5. The flow is compared on the basis of flow interaction between shear layer at anterior edge of cavity and the leading-edge shock generated at nose of the missile body. As flow speed is decreased from Mach 5 to Mach 2, this flow interaction has minimal impact. In this paper, two-dimensional CFD simulations in ANSYS FLUENT are carried out on the missile body at an angle of attack of 0 degree 9and Mach number 2.00 and 5.00 with density-based solver, assuming the ideal gas conditions with k-ℰ turbulence model. The study indicates that with decrease in flow Mach number, the flow characteristics changes between cavity and upper surface of missile. The plot of pressure coefficient distribution over missile surface shows significant variation in flow between cavity and upper missile surface with change in freestream Mach number. Configuration moving at supersonic speed of Mach 2 has less concern with the flow interaction and is rather affected by the detached bow shock wave with higher flow turn angle. Due to the compressible behaviour of the flow, a Prandtl-Meyer expansion fan is observed at shoulder of missile nose tip and also at rear end of missile. The effect of flow interaction between the shear layer at front edge of cavity and system of shock wave is major contributor for change in aerodynamic characteristics for model moving at speed of Mach 5. While for Mach 2, strong detached bow shock wave at leading edge of missile is the primary reason.