In supersonic water vapor flows of low pressure turbines, the nucleation phenomena and consequent condensation is commonly observed. Internal heat transfer which is caused by phase change is strongly irreversible and has unwanted effects on turbine efficiency. Also, the strike of formed droplets on the surfaces results in large amounts of mechanical damages. Condensation heat release to supersonic flow is named condensation shock and leads to a considerable pressure rise which in turn reduces the outlet velocity and occasionally causes severe oscillations and makes the flow supercritical. The authors have presented a novel analytical approach for the reduction of these unwanted results in Laval nozzles by volumetric heating of the convergent section. In this paper and in continuation of the series of papers by the authors, one dimensional, supersonic and two-phase flow is modeled analytically and the simultaneous effects of volumetric heat transfer and friction in the convergent nozzle are investigated. It is concluded that the simultaneous use of friction and volumetric heatingcan be an appropriate and useful technique for the control of two-phase flow conditions and keeping them in the desired range.

Vulnerability mitigation of war planes against modern sensor threats is implemented by reducing their physical signatures. This measure, as an inherent camouflage, is one of the important principles of passive defense and is emphasized in field of stealth technology of military platforms. In this essay, the changing effect of geometry of aircraft exhaust nozzle on the rapid cooling of ignition product released in the atmosphere to suppress IR signature has been put on high interest. The plume IR signature is investigated and to estimate its radiation rate, the FLUENT software is used. In this regard, the plume flow field for two convergent nozzles with different geometries was simulated with this software and geometry effects on plume temperature distribution was observed. The simulation result shows that the temperature of the hot plume core, for plane and notched nozzle respectively about 900° and 840° degrees Kelvin is obtained. Result of this research compared with the experimental results, was satisfactory and shows that the plume cools faster in a notched nozzle.

In this study، convergent nozzle ejector in the PEM fuel cell system is analyzed. This method can reduce the parasitic power of the fuel cell، recycle the unconsumed hydrogen to the fuel cell to increase the fuel usage efficiency، utilize the pressure potential energy of hydrogen and regulate the anode humidity with the recycle gas. For this purpose، continuity، momentum، energy and state equations are solved by numerical methods and effects of pressure drop (through the channel towards the anode)، operating pressure and temperature of the fuel cell and also nozzle diameter on the ejector performance was analyzed. With decreasing of pressure drop، even in primary lower pressure، increasing of performance pressure the performance of ejector will improved. The temperature increase has no effect on the performance of the ejector itself، but has enormous effect on the fuel cell. Increasing the diameter ratio of the constant diameter zone to the nozzle diameter leads to increasing of recirculation anode line of the fuel in higher pressure.

An axisymmetric jet issued from a sharp edged circular convergent nozzle has been studied experimentally in the near field at Reynolds number 5.3x104 to search the saddle shape behavior of mean velocity and the role of mean static pressure in the jet flow. Due to the vena contracta effect the centerline mean velocity of the jet increases about 4.2% of its exit value and corresponding decrease of mean static pressure is found in the potential core of the jet. Streamwise mean velocity shows saddle shape profile adjacent to the mixing layer near the exit (due to secondary current). Velocity peaks are associated with corresponding drop of mean static pressure. This saddle shape behavior is found to die down at the end of the potential core. The geometric virtual origin is at downstream of the nozzle tip and the entrainment rate is increased nonlinearly with the axial distance. A partial self-preservation of mean velocity is attained earlier than that of mean static pressure.

Heat transfer plays a major role in the design and performance of missiles and vehicles that have rocket driving. The main importance of this discussion is related to the safety range missile materials, especially in areas such as the combustion chamber and nozzle, which are under critical heat. Only one failure is enough to interfere with the operation of missiles. It may also be due to excessive heat in the nozzle throat, a portion of the sheet metal locally damaged. In which case the exhaust gas flow becomes asymmetrically, which led to the creation of inappropriate force vector direction. This leads to an error in the missile guidance to the target and a lot of fuel consumed on the spacecraft to correct its path by secondary driving system. In this project the amount of heat transferred to the chamber for different sections of engine during the simulation by software is calculated, also the coolant conditions and temperatures have been calculated in accordance with existing heat and cooling operation performed in each section. For this purpose, the engine divided into four parts, the combustion chamber, the convergent nozzle, throat nozzle and the divergent nozzle. The rate of heat transferred from the gas chamber (according to the conditions of gas in different sections) as well as the chamber temperature in the vicinity of the coolant and coolant temperature calculated for different sections of each section. Notably, in this research, ANSYS software package is used to create geometry, mesh and simulations.

Gas aerothermodynamics is the thermodynamics of a gas in high velocity associated with heat transfer. One of the devices which takes advantage of this field is a rocket system. High velocity flow filed with intensely varying pressure and temperature along the nozzle axis leads to reduction of residence time. This means the balance between chemical and flow time scales is changing in the flow stream. In this study, reacting flow composition variations in combustion chamber and nozzle of a liquid rocket engine is numerically investigated for hydrogen and kerosene fuels regarding combustion efficiency and performance of the engine. Numerical modeling has been conducted with the aid of commercial code FLUENT. k −e turbulence and EDC combustion models have been used to consider turbulence effects on the flow field. 9-step and 14-step skeletal mechanisms have been utilized to represent chemical oxidation of kerosene and hydrogen, respectively. Results show a reasonable accuracy in comparison with experimental measurements. As a result of this study, it can be concluded that in combustion chamber and convergent nozzle, the frozen and finite rate modeling have almost same results in all equivalence ratios. However, in divergent section of the nozzle some reactions proceed in finite rate regarding to fuel type. Finally, it can be noted that maximum Damkohler number occurs in the same axial position for both kerosene and hydrogen fuels.

This study experimentally explores the effect of tabs with asymmetric projections on the mixing effectiveness of jets at different nozzle exit Mach numbers with subsonic ranges of 0.4, 0.6, and 0.8. The results obtained with the tab-controlled jet are compared with those of uncontrolled jets. In this experimental investigation, a pair of identical tabs is deployed along a diameter of a convergent nozzle with inlet and exit cross sections of a circle, where each tab has two triangular projections configured at locations offset to each other at a distance of 1 mm on a plain rectangular stem. The geometrical blockage due to the presence of both tabs is maintained at 5.09% to minimize the thrust loss incurred due to tabs. The counter-rotating vortices generated at different locations of the tabs, caused instability or shear distortions at the nozzle exit, promoting jet mixing and eventually leading to rapid velocity decay along the jet axis and accentuating the reduction of the potential core. Compared to plain jet, reductions in core length of about 70%, 76%, and 81% at Mach 0.4, 0.6, and 0.8, respectively, are observed with the tab-controlled jets. The total pressure decay characteristics in the radial profile along the tab and normal-to-tab orientations have shown significant distortion in the jet structure, making it asymmetrical again owing to the asymmetrical positioning of projections on the tabs. Besides, in comparison with the plain nozzle, the total pressure decay characteristics in the radial profiles of tab-controlled jets are significantly different along the axial locations in the downstream direction due to the same reason of the asymmetrical positioning of triangular projections on the tabs. The primary research goal of this experimental investigation with asymmetrical tabs is to promote jet mixing asymmetrically to achieve thrust vectoring of jets.

This work simulates the turbulent boundary layer of an incompressible viscous swirling flow through a conical chamber. To model the pressure gradient normal to the wall, the radial and tangential velocity components across the boundary layer have been calculated by both the integral and numerical methods. The numerical solution is accomplished by finite difference, based on the finite volume method. The results show that the radial and tangential boundary layer thicknesses depend on the velocity ratios, Reynolds number and nozzle angle. The peak of radial and tangential boundary layer thicknesses are located at z/L≈0.2 and z/L≈0.8 from the nozzle inlet, respectively. Due to the short length of the nozzle, the contribution of momentum change on pressure loss is more significant than that on the shear stress. Also, the pressure gradient normal to the wall had been considered more accurately than that of the previous attempts.

In this study, a mathematical analysis by considering the effect of an actual index of expansion clearly shows a persistent of the existence of subsonic flow after the throat to a down stream in the region of divergent part to produce a supersonic velocity at the exit of a convergent-divergent nozzle. The length of the divergent part where subsonic velocity found is dependent upon the magnitude of the nozzle efficiency and the actual index of expansion. The change in velocity from subsonic, sonic and supersonic occurs only in the divergent part while the corresponding frictionless behavior has the classical features (subsonic in the convergent, sonic at the throat, and supersonic in the divergent). This is mainly due to thermodynamic processes which result a change in enthalpy due to friction and a gain in entropy. The reference conditions are newly derived for an actual frictional flow condition. This design aspect differs in a physical manner corresponding to that from an isentropic flow. An actual nozzle shape for convergent-divergent nozzles is also investigated under a non-isentropic flow condition.

The nozzle, an end-element of the propulsive process Cycle, represents a critical part of any aerospace vehicle. The task of accelerating and efficiently exhausting combusted and reactive gases according to the delivered thrust represents the main objective of the propulsion system design. Flow separation in supersonic convergent–divergent nozzles has been the subject of several experimental and numerical studies in the past. Now, with the renewed interest in supersonic flights and space vehicles, the subject has become increasingly important, especially for aerospace applications (rockets, missiles, supersonic aircrafts, etc). Flow separation in supersonic nozzles is a basic fluid dynamics phenomenon that occurs at a certain pressure ratio of chamber to ambient pressure, resulting in shock formation and shock/turbulent-boundary layer interaction inside the nozzle. From purely gas-dynamics point of view, this problem involves basic structure of shock interactions with separation shock, which consists of incident shock, Mach reflections, reflected shock, triple point and slip lines. In this article A Review on Flow Separation Phenomenon for Supersonic Convergent–Divergent Nozzles has been investigated.