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In this article, a mixed flow pump was designed using CFturbo software and analyzed during stable and transient process using CFX software. For the accuracy of the results of numerical analysis, a validation was done, and the numerical results were in good agreement with the experimental results. During the transient process in the numerical analysis, an instantaneous maximum pressure was observed. Pump head changes in this case are divided into three steps. In the first step, the head increases linearly, then in the second step, a momentary maximum pressure occurs, in the last step, the head curve tends to be stable. In addition, a parametric study was carried out in steady and transient mode, which includes the investigation of the effect of changing the blade outlet angle in four different angles in steady mode and two angles in transient mode, as well as the effect of changing the pump speed on its performance. By changing the speed of the pump impeller, the head and efficiency also increase or decrease in proportion to the increase or decrease of the speed. By changing the blade outlet angle, it was observed that with the increase of the outlet angle, the efficiency first increases and then decreases after the angle of 26 degrees, also the pump head increased by about 23 percent.

Increasing the propagation speed of the flame due to centrifugal force can lead to reducing the length of the combustion chamber and increasing the thrust-to-weight ratio in air gas turbine engines. The effect of centrifugal force on the propagation of the premixed flame has been investigated. For this purpose, the large eddy simulation of premixed combustion of the air-propane mixture in two straight and curved ducts with a step in the outer wall as a flame holder was performed using OpenFoam software and compared with the experimental data. The ducts have an inlet and outlet, and the averaged temperature and the wrinkling (the ratio of laminar to turbulent flame speed) were investigated for two different inlet velocities. It was observed that the curved duct inducing centrifugal force to the fuel and air mixture causes better mixing and wrinkling, increases the area of the flame, and as a result, the speed of flame propagation was improved. Also, the curved duct can withstand increasing the inlet velocity to higher values. To study the effect of fluid circulation, a new duct geometry for more mixture circulation was designed and analyzed. The comparison of temperature and wrinkling parameters in the outlet section for two initial curve ducts (C2) and the new one (C3) showed that the increase in the rotation due to the increase in centrifugal force improved the average temperature and wrinkling parameters in the outlet.

Various methods are used to control and attenuate the adverse effects of shock wave/boundary layer interactions (SWBLIs) and improve the performance of supersonic aircraft inlet. In this study, passive slot control is presented as a new, practical method for this purpose and its effect on supersonic inlet performance is investigated. To do so, a rectangular supersonic air inlet is first modeled and meshed at design Mach number of 2.2 and mesh independency is examined using three-dimensional (3D) computational fluid dynamics (CFD) method to achieve an ideal mesh; the most suitable mesh is selected to continue the work. The results of numerical simulations are compared with those of other studies and validated to ensure the correctness of the solution. In this paper, the inlet performance is first discussed without slots at design Mach number of 2.2 and, then, its performance is evaluated by creating slots and considering various parameters. The results indicated that using slots increased the pressure recovery and, consequently, enhanced the inlet performance in off-design conditions. It was also found that creating slots caused the vertical waves to reach the inlet with a delay and effectively controlled the flow. Then, after selecting the best slot, the inlet performance was presented and compared in off-design conditions at Mach numbers of 2 and 2.4, both with and without slots; finally, the best geometry was determined

Increasing the flame propagation speed with aid of centrifugal force proposed by Lewis is a new challenge that can reduce the length of the combustion chamber and thus increase the thrust to weight ratio. In this study, large eddy simulation of premixed combustion of air-propane mixture in a two-dimensional tube with closed ends have been implemented in OpenFoam Software to investigate the effect of centrifugal force on the flame propagation speed. Comparison of numerical solution results with experimental data showed about 8% error in the most critical centrifugal acceleration (3000g). To investigate the effect of the turbulence model, the combustion simulation using the k turbulence model for 3000g, was compared with the LES model and it was observed that the LES model, investigate propagation speed and flame wrinkling with higher accuracy. Considering the flame surface wrinkling parameter, it was observed that at a centrifugal acceleration equal to 4000g, the flame surface wrinkling, suddenly increased and then decreased rapidly that indicates the extinction of the flame. Then, the effect of pipe length on flame propagation speed and flame wrinkling at 2000g was investigated and it was observed that increasing pipe length and distance from the origin of rotation, the induced centrifugal force increased and as a result, propagation speed and wrinkling Increases.

Air intakes play an important role in the operation of aircrafts, so their optimal performance can have a significant effect on the performance of the propulsion system. The purpose of this study was to design and optimize an axisymmetric air intake for a mass flow rate of 10 kg / sec in a free-stream Mach number of 2.5 at sea-level conditions. In the present study, pressure recovery and flow distortion coefficients are selected as functional parameters for optimization. Initially, the parametric design of the intake and the selection of geometric parameters were dealt with, and then the interval of the parameter changes was determined. In this research, the NSGA-II multi-objective genetic algorithm was used as an optimization algorithm; artificial neural network was used to predict intake performance in the optimization loop; for training neural networks 243 Initial geometry has been designed and numerically solved. The genetic algorithm that used has 20 population per generation and 1000 generations. After 1000 generations, the resulting population is selected as optimal geometries. At the end of the optimization, the pressure recovery and flow distortion were improved by 4.4% and 49%, which indicates the efficiency of the optimization process.

The ability to reproduce a given distribution of total pressure loss by horizontal arrangement of splitting airfoils has been evaluated numerically in this research. This array arrangement is a proposal design for a new distortion generator mechanism. Due to significant differences between steady-state simulation based on RANS turbulence models and reference experimental test results, unsteady simulation was performed with vortex shedding frequency- proportional time step. The 90-degree airfoil reference test measurements show that maximum loss zones from primary downstream stations to fully developed wake region switch from vertical placement to horizontal; this distinctive feature was not seen in the unsteady RANS results. Detailed simulation of downstream turbulent mixing based on DDES hybrid model, predicts above switching well. The repeatability studies of single splitting airfoil loss pattern in airfoils horizontal arrangement combined pattern, led to necessity of 90-degree airfoil downstream flow structure correction that is a destructive factor of pattern predictability. This correction can be achieved by changing the airfoil aspect ratio.

Inlet in ramjet engines are used to reduce the air flow from free stream to the required speed for flame stability in the combustion chamber. Hence the precise and correct design of inlet have rightly influence on whole system of ramjet engine performance. At the present work, an axisymmetric supersonic inlet with external compression has been designed based on requirements of the ramjet engine at Mach 2.1. Steady flow simulation was done for inlet individually by CFD software in different fuel flow and Mach numbers to investigate the inlet performance that integrated by other parts of engine. Then transient simulation performed for integrated engine by selection of a special fuel flow-time function that injected in combustor. The results of simulation exhibited the shock velocity and performance of inlet that integrated by the engine. The results show that the designed geometry can meet the desired performance requirements considering the fuel control system ability.

Grain design is the most important part of a solid rocket motor. The aim of this study is finocyl grain design based on predetermined objective function with respect to ballistic curves in order to satisfy various thrust performance requirements through an innovative design approach using a genetic algorithm optimization method. The classical sampling method has been used for design space-filling. The level set method has been used for simulating the evolution of the burning surface in the propellant grain. An algorithm has been developed beside the level set code that prepares the initial grain configuration using Pro/Engineer software to export generated models to level set code. The lumped method has been used to perform internal ballistic analysis. Two meta-models are used to surrogate the level set method in the optimization design loop. The first method is based on adaptive basis function construction and the second method is based on the artificial neural network. In order to validate the proposed algorithm, a grain finosyl sample has been investigated. The results show that both grain design method reduced the design time significantly and this algorithm can be used in designing of any grain configuration. In addition, data have more accuracy in grain design based on the artificial neural network, so this method is the more effective and practical method to grain burn-back training.

Grain design is the most important part of solid rocket motor design. In this paper, the goal is implementing and comparing various Artificial Neural Network metamodels in Finocyl grain design based on predetermined objective function with respect to thrust history or pressure history in order to satisfy various thrust performance requirements through an innovative design approach using genetic algorithm optimization method. The classical sampling method is used for design space filling. The level set method is used for simulating the evolution of burning surface in the propellant grain. An algorithm is developed beside the level set code that prepares the initial grain configuration using Pro/Engineer to export generated models to level set code. The zero dimensional method with considering the effect of erosive burning is used to perform internal ballistic analysis. The meta-model is used to surrogate the level set method in optimization design loop. The surrogate method is based on artificial neural network based on Multi Layer Perceptron with comparing various training algorithm and Radial Basis Function. Finally, a case study is done to verify the proposed algorithm. Observed results show that grain design method reduced the design time significantly and this algorithm can be used to design of any grain configuration.

To analyze the behavior of fluids around obstacles cylinder to reduce costs, reduce the time to achieve the results and analysis of boundary conditions to the numerical simulation using FLUENT software Rvym.dr paper, Unsteady Flow, two-dimensional, incompressible intended for each Reynolds static pressure contours, Vorticity, flowing lines and speed has been obtained.It can be seen that the rectangular cylinder pressure at the inlet of the circular cylinders and more triangular and wherever is made of vortices And in which the pressure Darym.dr least the low pressure flow lines will be able to lose your makeup.Also generally at high Reynolds numbers for all cases, and in this case also during the more turbulent flow vortices more.The main gates of the high-pressure occur at these points there is no vortices and flow irregularities.It is known that the contours pressure cylinder with a rectangular cross-sectional area drag to endure more.

The aim of this paper is the optimum design of a single circumferential groove for NASA Rotor 37, a transonic compressor. The design variables are the width, the depth and the position of the groove. Also the objective functions are the stall margin and the peak efficiency. The total pressure ratio at the near stall condition is used as a criterion to choose the optimum design. Neural networks are used to model the relation between design variables and objective functions. 123 various design points are simulated by CFD and the results are used to train the networks. Afterwards, a multi-objective genetic algorithm is used to optimize the design based on the maximum stall margin and the minimum reduction of the peak efficiency. After sorting the Pareto front according to the stall margin value, the first design point whose total pressure ratio is larger than the smooth wall condition, is selected as the optimum design. Simulation of the compressor with optimum groove shows the increase of the stall margin by 6.2 percent with negligible effect on the efficiency.

In designation process of liquid rocket engine, selection of cooling rout of thrust chamber have important role for assessing safety and reliability of the engine. Also the correct perception of internal flow characteristics in LRE thrust chamber, such as temperature and pressure, plays effective role in optimization of engine design to decrease consumption fuel and how to cooling the thrust chamber wall. One of the ways to cooling the thrust chamber wall to decrease the combustion temperature near the wall is injection extra fuel in this section. In this paper, numerical investigation of the effect of film cooling with mixed heat transfer on decrease the thrust chamber wall temperature in the LRE, relative to condition that only exist mixed heat transfer in the wall was performed.

Because real experiments on a large scale is costly, And all the different scenarios will not gain experience, We go to the numerical simulation. In the present study has been simulation of flow in a channel with six cylindrical barrier. Fluid flow parameters Has been obtained Around the cylinder to two different arrangement. Also has been obtained at different Reynolds numbers for both, fixed cylinder and rotating cylinder. So that both laminar and turbulent flow in the covers. In this article has been obtained by detailed simulation and validation results, contours speed, Vorticity, pressure and flow lines using Fluent software. It can be seen that lower drag coefficients into the cylinder with increasing Reynolds number. As well as rotating cylinders experience less drag than the fixed cylinders. The results show that the arrangement of cylinders affect on the flow parameters. Finally, relying on the results will suggest a suitable arrangement for the construction of piers and offshore platforms.

The present research is mainly devoted to investigate the effect of pressure ratio (ᴨ) on the Shock Wave-Boundary Layer interaction phenomena along the wall surface in internal compressible flows. The implicit Reynolds Average Navier-Stokes equations and k-ω turbulent model have been used to numerical simulation and investigation the effect of pressure ratio on location, structure of the shock train and the boundary layer specification. The obtained results show that the present numerical results are in good agreement with the experimental data. On increasing ᴨ, not only the shock train moved toward the upstream flow but also the shock structure changed from X-shape to λ. Furthermore with increasing ᴨ the displacement of shock train along the duct is increased. With decreasing ᴨ the shock train moved to the downstream so that the length of the duct isn’t enough for complete recovery pressure. The flow separation bubble is formed near the wall, when the shock was strong. The experimental results of supersonic blow down wind tunnel in Cambridge University have been used for validation.

The flow fields over a non-slender lambda shaped wing including vortex formation and the effect of angle of attack on structure, strength and location of the vortices were investigated in a closed circuit low speed wind tunnel by pressure measurement. The tests were conducted at velocity of 20 m/s and different angles of attack from 5 to 20 degrees. The Reynolds number of the model was about 2×105 according to the root chord. The results showed three vortical flows were formed over the wing surface. The first vortex was the apex vortex; the second one was generated before the leading edge crank and strengthened after the crank and the third one was formed after the leading edge crank. By increasing the angle of attack the vortices became larger and stronger and closer to the wing root. At a specific angle of attack the structure of the vortices was changed rapidly and the vortex break down was occurred. The location of vortex break down moved toward the wing apex by increasing the angle of attack.

The aerodynamic coefficients characteristics over a lambda-shaped flying wing aircraft with 55°-30° leading edge sweep angles have been investigated in a closed circuit low speed wind tunnel. The experiments were conducted at tunnel velocity of 90 m/s, the angles of attack of -6 to 17 and the side-slip angles of -8 to 8 degrees. All forces and moments were measured using an external six-component force balance located below the wind tunnel. The wall corrections were also performed for all test conditions. To improve the aircraft longitudinal stability characteristics, a new model with an increased leading edge sweep angle of 2 degrees were also tested and compared with the original model. A “pitch-up” phenomenon identified to occur at a rather low angle of attack of α=7.7 degrees, although it occurred at the higher angle of attack of α=8.7 degrees for the increased swept angle model which means an increase in useable lift of the aircraft. Moreover, off-surface pressure measurement over the wing surface was conducted to examine the onset and development of the flow separation over the wing surface. The results showed that the flow separation started at the trailing edge crank location and extended to the other parts of the wing, especially the outer wing.

In this research is prepared the optimum design muffler code for maximum silence in perforated muffler to use in APU exhaust. According to far Annex 16 the power of sound (SPL) must be less than 90 dB on ground from aircraft. One of the sources of noise on ground is the exhaust noise of APU. The APU is used for product pressurize air for main engine start and air condition and for produce electric energy. So the muffler must be use in APU exhaust. In this paper was studied the method of calculation and effective parameters on transmission loss. After that, a code was produced for design muffler with respect to limitation of dimension and designer requirement. For verification of design code, the result of this code was compared with test result. By this code the muffler was designed for respective aircraft by attention to limitation of situation.

In this study, the vertical flow on a sharp edged, 70 degrees swept back delta wing was experimentally investigated in a smoke tunnel, using laser sheet technique. Previous studies show that changing Reynolds number has little effect on the vortex structure of sharp edged delta wings, although the angle of attack has the major effect. Furthermore, Spalart Almaras model is used for numerical investigation on the delta wing. The effect of angle of attack on size and break-down location of the vortices on the wing was studied. The results show that increasing of the angle of attack increases the size of the vortices and the height of the vortex core to the wing surface as well. The bigger vortex on a delta wing leads to increasing the lift of the wing and it increases the maneuverability of the aircraft. Abrupt changes occur in the structure of the vortices at very high angles of attack which is designated as vortex break-down. The vortex break-down cause’s intense oscillation of the surface pressure of the wing and it decreases the aircraft maneuverability.

In order to perform experiments in wind tunnels it is needed to arrive in the “run” stage, and knowing that the “run” stage is arrived after passing the triggering and “start” stage, it is necessary to study the tunnel flow in the start stage. In the start stage, which is occurring in a very short time interval in an unsteady and transient mode, some strong shock waves are formed who should rapidly pass the test section to be swallowed by the second throat of the supersonic diffuser. Otherwise, the normal shock waves formed in the test section will remain and in addition to causing severe pressure drop, will hinder the flow in the tunnel. Such a phenomenon will practically make the tunnel unusable. In this paper, the “start” stage of a blow-down supersonic wind tunnel with a Mach number of 2.8 and a nuzzle geometry designed with “method of characteristics”, has been studied under unsteady and transient conditions. Mach counters and their diagrams were plotted to assess the propagation of shock waves and achieve goal Mach number. y+ diagram was used to cross examine the modeling of the flow turbulence done in Spalart – Allmaras method. The results confirmed the accuracy of the model.

The manner of spray the fuel has an important effect on stability and convenient mixture of fuel and oxidizer. Swirl injectors because of formation of uniform spray cone and proper distribution of the fuel have been used in many rocket engines at aerospace industries. Formation and development of air core in the interior stream of swirl injectors and simulation of them is complicated. Double base-swirl injectors have better atomization and have more effective and uniform mixture and combustion. Inner stream has an important effect on spray characteristics. In this article we want to simulate liquid film، air core and spray angle in double base swirl injector. Inner stream of swirl injector simulate with two-phase method and we use air for gas-phase and water for liquid-phase and validate our results with experimental results. Because this way is very cheaper and faster rather than experimental studies.

In this paper, numerical simulation of cavitating flow over hydrofoils is presented by the aid of Preconditioned Artificial compressibility Method for solving the multi-phase N-S equations. Dominant equations include conservation of mass, momentum and a liquid phase volume fraction transfer equation. Equation of volume fraction transfer is obtained on the basis of homogeneous equilibrium flow concept in modeling of multi-phase flows. Mass transfer between phases is simulated by the Merkle model. To accelerate the convergence rate, the discrete equations are preconditioned. Third order Roe-based flux difference splitting and second order central differencing method respectively is utilized for convection and viscous term discretization. Effects of turbulence are simulated in terms of an eddy viscosity coefficient added to molecular coefficient of viscosity via Spalart-Allmaras model. To demonstrate the capabilities of the scheme, several single and two-phase test case problems are computed and the results are compared with experimental and numerical data. Computed results present the appropriate accuracy of algorithm. Results of single phase flow over NACA0012 hydrofoil at renolds number, 2.8e6 & 0, 6 angle of attack, are compared with experimental data. The capability of code for simulation of cavitating flows at various cavitation numbers and angle of attacks are presented via simulation of two phase flow over non-truncated & truncated modified NACA0009 hydrofoil at renolds number, 2.0 e6 & 0, 2.5 angle of attack & σ = 1.0, 0.9, 0.8, 0.75, 0.6, 0.5, 0.4, 0.3. Then results are compared with experimental and numerical data. The results are in good agreement with the available data and other published computations.