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The mixing enhancement and core length reduction of a jet without significant loss of thrust are essential for reducing infrared radiation, mitigating aeroacoustic noise, improving combustion characteristics, and thrust vectoring. The jet mixing can be improved by manipulating the flow behavior. In subsonic and sonic jets, the flow manipulation may be achieved by utilizing nozzles with non-circular geometries that shed vortices of varying size due to their non-uniform azimuth curvatures. Non-uniform vortices generate differential spreading along the nozzle’s perimeter, causing axis switching and improving entrainment characteristics. Therefore, the present study examines the effects of two non-circular nozzle exit shapes (elliptic and square) on the mixing augmenting efficacy at subsonic and sonic flow conditions. The circular nozzle is tested for comparison. Both quantitative and qualitative analyses evaluate the efficacy of nozzles with non-circular exit geometries. Among the configurations investigated, the elliptic nozzle is superior in shortening the potential core length and enhancing the jet spread. A maximum reduction of 18.75% in core length with rapid jet decay was accomplished with the elliptic nozzle. The measurement of pressure profiles at different streamwise locations reveals that the spread rate is greater for elliptic and square jets than their circular counterpart. The elliptic jet exhibits the highest spread along the minor-axis direction compared to the major-axis direction. The differential jet spread rate in the elliptical jet causes an early axis-switching––direct evidence of mixing augmentation. Shadowgraph images show the asymmetric pattern of shock cell structures and differential spreading in elliptic and square jets.

The primary focus of the present survey is to categorize the results of various investigations on the Shock/Boundary-Layer Interactions (SBLIs), their repercussions, and the effective ways to control them. The interactions of shock waves with the boundary layer are an important area of research due to their ubiquity in several applications ranging from transonic to hypersonic flows. Therefore, there is a need for a detailed inspection to understand the phenomena to predict its characteristics with certain accuracy. Considering this in mind, this article presents some key features of the physical nature of SBLIs, their consequences, and the control techniques in a sequential manner; in particular, the passive control techniques for the supersonic and hypersonic intakes are reviewed in detail.

The desired changes in flow characteristics are obtained by flow control, which implies manipulating flow behavior such as drag reduction, mixing augmentation, or noise attenuation, employing active or passive devices [...]

The length of the supersonic jet ejected from the military aircraft must be reduced in order to decrease its heat signature and aeroacoustic noise and thereby to enhance its stealth capability. The reduction or manipulation of the supersonic core can be achieved through various passive control techniques. Considering this, the present study explores the mixing characteristics of supersonic jets with and without passive controls. Passive controls in the form of grooves configured at the exit of a Mach 1.73 convergent–divergent nozzle are investigated computationally. Particularly, the supersonic jet decay characteristics and flow development for a plain nozzle and a nozzle with semi-circular, square, and triangular grooves are presented. In addition, the study explores different turbulence models, namely, Spalart–Allmaras, realizable k-ε, std k-ω, shear stress transport (SST) k-ω, and SST transition. The realizable k-ε turbulence model is found to be the most effective one in capturing the supersonic jet structure. It is observed that the grooves produce large distortions in the jet structure, accompanied by significant mass entrainment and lateral spread. Interestingly, semi-circular grooves are proven to be most effective in all cases of expansion level than square and triangular grooves. For the semi-circular grooves, a maximum of 48.5% reduction in the supersonic core length of the correctly expanded jet at nozzle pressure ratio (NPR) of 5 is achieved. The reduction in the supersonic core length for semi-circular grooves is 31% for the overexpanded jet at NPR 4 and 29% for the underexpanded jet at NPR 7.

The present experimental investigation explores the ability of the corrugated actuator-tabs (or simply, the corrugated actuators) of aspect ratio 2 in encouraging the mixing of Mach 1.73 circular jet at different levels of expansion conditions. Essentially, rectangular, triangular, and semi-circular corrugated actuators are deployed at the opposite positions across the diameter of the nozzle outlet and compared with the plain or simple actuators located at the same positions. Quantitative and qualitative observations were conducted by pitot pressure measurements and shadowgraph flow visualization. Interestingly, the characteristic decays are higher for the semi-circular and triangular corrugated actuators, exhibiting superior mixing than the rectangular corrugated actuator in the far-field. However, the core length reduction for rectangular corrugation is greater than for triangular and semicircular corrugation geometry. As high as 96.5 % core length reduction was obtained for the rectangular corrugated actuators operating at the overexpansion condition. The pressure profiles along and perpendicular to these actuators confirmed that the deployment of the corrugation geometry significantly reduces the flow asymmetry. The shadowgraphic flow visualizations reveal that the plain and corrugated actuators weaken the waves and shorten the cells when compared to the uncontrolled or free jet.

The accumulation of waste in one place for much time makes it rotten and which leads to the attraction of flies and rodents, which can harbor many of health hazardous conditions? Likewise, some have the cure as burning waste, and this is worst planned for getting rid of this problem of waste as this further it pollutes air so as living being to as we inhale it. To avoid such disastrous situations, we are going to make these smart bins named as "IoT enabled Devices for Waste Management'. These dustbins are interfaced with IOT system with MSP-EXP430G2 Launchpad using Wi-Fi and GPRS modem. Hence, the state will be updated on to the HTML page along with it we have an App it will be making this system to be user handy for everyone. The heart of this system is a Wi-Fi module and GPRS modem; essential for its real-time data updates for picking up the team to function fluently. The main aim of this project is to protect the environment, peoples as well as the proper management of resource by adequate waste management devices.

The neural network–based aircraft parameter estimation techniques have gained prominence in the last decade. Neuro–Gauss–Newton (NGN) technique is a widely used neural network–based algorithm. Although the NGN technique is useful for aircraft parameter estimation, it has some limitations. The present study is motivated by the limitations of the NGN method in estimating aircraft parameters, and hence, a new hybrid Luus–Jaakola/Hooke–Jeeves (LJ/HJ) method is proposed. The hybrid LJ/HJ method is based on two direct search methods: the Luus–Jaakola (LJ) method and the Hooke–Jeeves (HJ) method. In the hybrid LJ/HJ method, a global search of the minima of the error cost function is performed using the Luus–Jaakola method in the first phase. After that, the final solution is obtained using the Hooke–Jeeves method. The parameters estimated by the hybrid LJ/HJ method are compared with those estimated using the NGN method. The results obtained from the hybrid method are accurate, and unlike gradient-based optimization methods, the convergence issues are not observed during the estimation of parameters.

To have an efficient combustion, the aircraft engine intakes operating at supersonic and hypersonic speeds necessarily decelerate the flow to subsonic level before entering the combustor, which is achieved by a combination of oblique and normal shock waves in the intake-isolator, and thus the high-speed intakes are called the mixed-compression intakes. The advantages of shock-enabled compression, however, does not come standalone rather associated with colossal losses due to shock and boundary layer interactions (SBLIs). The repercussions in the flow due to these interactions may include; intake unstart, abrupt thickening, separation of the boundary layer, unsteady shock oscillations, etc. Therefore, the SBLIs must be controlled to minimize the losses. Control of these interactions by manipulating the boundary layer using micro-vortex generators (MVGs) has gained prominence. In this study, a new ramped-vane MVG configuration, deployed near the shock impact point in the Mach 2.2 mixed-compression intake at varied contraction ratios, has been experimentally investigated. Plain intake and the intake controlled with conventional MVGs are also investigated for comparison. The heights of all the MVGs were varied as; 600 μm, 400 μm and 200 μm. The ramped-vane MVGs of height 200 μm are found to be the most efficient in causing a favorable pressure drop at the locations; near-upstream (x = 0.48 L) and near-downstream (x = 0.7 L) of the MVGs. The maximum reductions in static pressures about 11% at the intake contraction ratio of 1.20 at x = 0.48 L, and about 24% at the contraction ratio of 1.23 at x = 0.7 L, are achieved. The Schlieren pictures clearly demonstrate the effectiveness of all the tested ramped-vane MVGs (particularly 200 μm MVGs) in weakening the waves and reducing the separation length.

The occurrence of shock and boundary layer interactions (SBLIs) in aircraft-engine intakes often results in poor total pressure recovery, a degraded boundary layer, and, in worst cases, flo...

The attitude control of a solid rocket engine through the control surfaces becomes difficult particularly when the rocket is large and the payload is high. In such cases, an obvious means of producing forces for controlling the flight is the deflection of exhaust gases, commonly known as the gas-dynamic steering or thrust vector control. In this study, the effect of a strut on the exhaust gas deflection, deployed in the divergent-portion of a Mach 1.84 nozzle, has been experimentally investigated. In order to evaluate the thrust vectoring efficiency of the strut at overexpanded, correctly expanded and underexpanded states of the jet, the settling chamber pressure is varied as 4 bar, 6 bar and 8 bar, respectively. The height of strut is varied as 1.5, 2.5 and 3.5 mm. The strut of height 3.5 mm was found to be the most efficient in causing the maximum jet deflection at the overexpanded conditions, with a maximum jet deflection of about 3.6°. These results are found to be in good agreement with the qualitative data.

Malicious software aka Malware is one of the stern threats that has been around ever since the computers were developed. Malware is designed by malicious attackers whose intension is to bring down the system or to take over the system. It is very difficult to detect unknown malware as compared to known ones. Steganography tools allow users for embedding the hidden data inside the carrier file like video, audio, or an image, and later data is extracted. Here, malware is hidden inside an image by attackers to manipulate the victims and it can be done via phishing technique which is one of the famous and well-known practices done by the attackers. This paper provides an overview of creating malware using Metasploit in the Parrot OS (Operating System) which is a secure OS, later the Steganography tool is used to manipulate the victim and the victim's system is taken over by the attacker. Also, a summarized analysis is given on the Open Source Intelligence Tools (OSINT) and which one gives the best results.

In aircraft engines operating at high Mach numbers, it is exigent to reduce the magnitude of flow speed from supersonic to subsonic before entering the burner, to accomplish a proficient ignition. It is accomplished by a progression of oblique shocks as well as a normal shock wave in supersonic intakes. However, the advantage of shock enabled compression in intakes does not come independent but with colossal losses due to shock-boundary layer interactions, which includes intake unstart and an abrupt thickening/separation of the boundary layer. Controlling these interactions by boundary layer control using micro-vortex generators (MVGs) has gained prominence. In the present study, an attempt is made to control the interactions at the shock impact point in a Mach 2.2 mixed compression intake. Two types of MVGs; a conventional configuration and an innovative ramped-vane configuration were experimentally investigated by varying the MVG heights as 600, 400, and 200 μm. Also, the near-field effects of MVGs are quantified in terms of static pressure variation in the internal flow. It is found that the innovative MVGs of height 200 μm leads to favorable pressure drop in both the upstream and downstream region, due to enhanced flow mixing near the boundary layer.

In order to achieve proficient combustion with the present technologies, the flow through an aircraft intake operating at supersonic and hypersonic Mach numbers must be decelerated to a low-subsonic level before entering the combustion chamber. High-speed intakes are generally designed to act as a flow compressor even in the absence of mechanical compressors. The reduction in flow velocity is essentially achieved by generating a series of oblique as well as normal shock waves in the external ramp region and also in the internal isolator region of the intake. Thus, these intakes are also referred to as mixed-compression intakes. Nevertheless, the benefits of shock-generated compression do not arise independently but with enormous losses because of the shockwave and boundary layer interactions (SBLIs). These interactions should be manipulated to minimize or alleviate the losses. In the present investigation a wall ventilation using a new cavity configuration (having a cross-section similar to a truncated rectangle with the top wall covered by a thin perforated surface is deployed underneath the cowl-shock impinging point of the Mach 2.2 mixed-compression intake. The intake is tested for four different contraction ratios of 1.16, 1.19, 1.22, and 1.25, with emphasis on the effect of porosity, which is varied at 10.6%, 15.7%, 18.8%, and 22.5%. The introduction of porosity on the surface covering the cavity has been proved to be beneficial in decreasing the wall static pressure substantially as compared to the plain intake. A maximum of approximately 24.2% in the reduction in pressure at the upstream proximal location of 0.48 L is achieved in the case of the wall-ventilated intake with 18.8% porosity, at the contraction ratio of 1.19. The Schlieren density field images confirm the efficacy of the 18.8% ventilation in stretching the shock trains and in decreasing the separation length. At the contraction ratios of 1.19, 1.22, and 1.25 (&lsquo, dual-mode&rsquo, contraction ratios), the controlled intakes with higher porosity reduce the pressure gradients across the shockwaves and thereby yields an &lsquo, intake-start&rsquo, condition. However, for the uncontrolled intake, the &lsquo, unstart&rsquo, condition emerges due to the formation of a normal shock at the cowl lip. Additionally, the cowl shock in the &lsquo, intake is shifted upstream because of higher downstream pressure.

In the present study, the characteristics of three-inline non-premixed Oxy-Methane turbulent flame jets, with Methane jet at the center and two Oxygen jets on either side, are computationally investigated. For all the jets, the velocity is varied from 10.13 m/s to 108 m/s. It is found that, in the presence of central jet, the mixing of lateral jets are delayed further downstream. In contrast, the central jet diffuses at a faster rate. At far-field locations, all the jets merge with each other and form a single jet, which can be seen from the uniformity of the radial velocity distribution. The turbulent intensity is found to be more at the jet periphery, where jets interact with the surrounding fluid. The temperature of the flame is found to be higher at the periphery of the methane and the oxygen jets, due to the existence of most appropriate equivalence ratio. Also, the flame lift-off height is found to be increasing with the jet velocity. The concentration of methane is reduced along the streamwise direction due to the penetration of combustion products towards the jet centerline.

In the past, many researchers have carried out water-hammer pressure analysis using Joukowsky equation. However, it has been observed that the computed pressure surge is no longer applicable based on the equation. The Joukowsky equation cannot be used even within the reflection time of the long pipeline. In such cases, the actual pressure rise due to the sudden closure of a quick acting valve will be several times more than that of the sudden increase in pressure as calculated by the Joukowsky equation. The phenomenon of rising pressure at the upstream of an instantaneously closed valve with the passage of time caused by the pipe friction is commonly called as linepacking. In this paper, various parameters affecting the linepack pressure have been thoroughly investigated. As the relative roughness increases, the resulting non-dimensional linepack pressure $$ \left( {P_{LP} /P_{o} } \right) $$ significantly increases and the proportionality constant was equal to 1.5. The linepack pressure was determined to be decreasing with increasing valve closure time. The dominant parameter that influences the linepack pressure is found to be the Reynolds number as compared to the Mach number, and the relative roughness. Furthermore, the linepack pressure is found to be proportional to frictional head loss $$ \left( {h_{L} /D} \right) $$ , and inversely proportional to inlet pressure $$ \left( {P_{o} /\left( {\gamma L_{o} } \right)} \right) $$ . Finally, a linear regression equation was developed in terms of non-dimensional variables to estimate the linepack pressure using hand calculations without undergoing numerical modeling procedures. The proposed equation was validated for sudden valve closure pressure histories available in the literature. The proposed method is applicable to long distance water supply pipelines where the linepack pressures are significant.

The present study numerically investigates some novel modifications to augment the performance of the H-type finned tube banks, which are widely used in waste heat recovery in industries. The imposed modifications upon the original H-type finned tube banks include the following: (1) Design 1 contains some triangular cuts at the edge of the original rectangular fin; (2) Design 2 modifies the original rectangular geometry into a trapezoid shape; (3) Design 3 renders the original rectangular cross-section fin thickness into trapezoid cross-section; and (4) Design 4 changes the original rectangular shape into a circular shape. Based on the simulations, it is found that Design 1 shows barely any improvements in the heat transfer performance and surface area reduction. Design 2 can provide some weight saving and surface area reduction at a slightly inferior heat transfer performance. Design 3 can offer up to 14% improvements in the overall heat transfer performance without any pumping power penalty. Yet, Design 4 provides the maximum weight saving as compared to the original reference case. With 3–9% lesser surface area than the reference case, Design 4 still offers marginally higher heat transfer performance.

The efficient combustion necessitates the flow speed reduction from supersonic to subsonic level before entering the burner, which is achieved by a series of oblique and normal shocks in the isolator region of the intakes. However, the advantages of shock-induced compression are compromised by the associated penalties due to shock-boundary layer interactions (SBLIs), leading to intake unstart and abrupt thickening/separation of boundary layer. The micro-vortex generators (MVGs) are well-accepted boundary layer control in subsiding SBLIs. In the present investigation, the Mach 2.2 mixed compression intake controlled using MVGs with varied heights has been experimentally studied. The MVGs of height 200 μm are found to be efficient in promoting mixing near the boundary layer with minimized interaction losses.

The interaction of a shock wave with boundary layer is called the Shock–Boundary Layer Interactions (SBLIs). This interaction has a profound effect on the performance of a vehicle flying at high Mach numbers, especially in hypersonic flow regime. Because of their ubiquitous presence at supersonic and hypersonic speeds and their influence on the vehicle performance, the studies on SBLIs have been a challenging task among the researchers for past seven decades. This chapter explores the fundamental aspects of SBLIs with an emphasis on understanding the physics behind these interactions. The major outcomes of a few experimental studies performed in the author’s laboratory are also discussed.

In an irrotational flow field, one can use either velocity potential or stream function to characterize the flow and both must fulfill Laplace equation. Besides, the investigation of this equation is substantially less demanding than the coordinate approach of completely viscous Navier–Stokes equations. The linearity of Laplace equation enables to add various basic solutions to obtain more complicated solutions. The investigation of these solutions of Laplace equation is referred to as potential flow theory, which has a great deal of practical applications in characterizing complex flow fields. In this chapter, we will discuss the velocity potential and stream function for a couple of rudimentary flow fields, for example, uniform flow, source/sink, doublet flow, and vortex; they can be superimposed to get complex flow fields of practical importance.

Thin airfoil theory is a straightforward hypothesis of airfoils that relates angle of attack to lift for an incompressible and inviscid flow past an airfoil. This theory idealizes the flow past an airfoil as two-dimensional stream around a thin airfoil which can be envisioned as tending to an airfoil of zero thickness and infinite wingspan.

The purpose of the present chapter is to review those aspects of thermodynamics that are essential to compressible flows. This is not the slightest bit expected to be a thorough discussion of thermodynamics; rather, it is an overview of only those key ideas and relations that will be of direct use in subsequent chapters. For the readers who are already familiar with the thermodynamics, this discussion will be a ready reminder of some important relations. This discussion is similarly helpful to those readers who are not comfortable with thermodynamics, as it is fairly independent in giving essential concepts and governing equations much of the time utilized as a part of subsequent chapters.