جستجوی مقالات مرتبط با کلیدواژه « امواج ضربه ای » در نشریات گروه « مکانیک »

In this study, a high-resolution scheme based on the WENO family has been developed in a pressure-based algorithm to prevent non-physical fluctuations based on Riemannian solver for steady and unsteady one dimensional and two dimensional compressible flows. The solution method is based on finite volume which uses an implicit solver with a structured collocated grid. Boundedness is applied by a high resolution essentially non-oscillatory scheme. To evaluate the numerical method, a shock tube containing shock wave, contact discontinuity and expansion waves has been considered, and the results obtained have been compared with the analytical results and the results based on the density based algorithm. The developed method is evaluated for lax configuration in two-dimensional inviscid flow. In addition, this method has been used to simulate two dimensional steady flow in a channel containing bump. The results show that the developed method is able to capture the physical and numerical discontinuities well.

In this study, shock wave mitigation technique was analyzed using qualitative observations of plasma discharge in Mach 2.45 and atmospheric conditions testing. Plasma was produced in front of the aero-spike model by a 50 Hz, 50 mA, 30Kv electrical discharge and shadowgraph imaging technique at 300 frame per second and camera recording at 1000fps were used to record the qualitative results. Laboratory results show that increasing the magnetic field increases the frequency, stabilizes the glow discharge, changes the motion path of the charged particles from circular to cyclotron, improves plasma overlapping and also thickens the shock layer. Shadowgraph images at Mach 2.45 show that combining magnetism with and increased by 7.5 degrees at a shock wave angle and mitigates shock waves and removes the bow shock downstream of the spike. This is the most important result that indicates combining plasma and magnetism can remove shock waves at supersonic speeds and thus reduce wave drag.

Shock waves are presented in hypersonic aircrafts. They increase drag and as a result of additional friction, surface heating increases. In this research, a wind tunnel model; a combination of a 60o slender physical spike, used as cathode and a 60o truncated cone- cylinder, as anode, were experimented in flows with Mach numbers 1.5, 1.95, and 2.45. Plasma was produced in front of the aero-spike model by electrical discharge of 50 HZ, 30 KVDC, and 50 mA. Shadow and plasma glow imaging techniques were used simultaneously for flow and plasma visualization. Shadow imaging, in the afore mentioned Mach numbers, shows that the plasma being discharged behind shock wave, in spite of increasing the magnetic field, has a slight effect on decreasing the intensity of the shock wave. With increasing Mach number, the Shock wave of the truncated conical nose moves downstream and as a result of the plasma discharge taking place below the nose and the constant magnetic field, the wave below the nose is eliminated. The experimental results indicate that at Mach number 2.45, the shock wave attaches to the truncated nose, thus; the continuous plasma discharge below the spike and in front of the wave eliminates the wave. This is the most important result of this study indicates that aero-spike plasma discharge can remove shock waves and thus reduce drag.

Simple designing، low maintenance and build cost، and high ratio of thrust to weight، also new methods of numerical simulating and solution causes to develop the Pulse Jet engine usage in non-military fields. Impact of geometrical parameters of valve less pulse jet engine on thrust has been studied in this article. First، study concentrate on main geometrical parameters like diameter and length، and transform those parameters to dimension less parameters (length to mean diameter ratio) for each section of the engine. Then the allowable range for parameters is achieved based upon experimental researches of Lockwood. Ten new geometries for valve less pulse jet engine has been defined and numerical solution of two-dimension flow field of inside engine has been presented with Fluent Code. New method has been presented (without solving the combustion) with considering combustion as initial condition. To ensure that the accuracy of solution is obtained، validation has been done with a valve pulse jet engine that shows excellent results with less than 5 percent error. Eventually the main result shows that amount of parameters that has most impact on thrust is 29 for exhaust pipe، 1. 25 for combustion chamber and 3. 5 for intake pipe. The volume of combustion chamber should not exceed the limits otherwise it will be caused non-uniform pressure distribution and will be effect on engine performance.

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.

An extensive experimental study has been conducted to investigate the performance and stability of a supersonic axisymmetric mixed compression air intake designed for a free stream Mach number of 2.0. Unstable flow conditions, where the self-sustained oscillations of the shocks waves occur, have been studied in this investigation. Aside from the buzz triggering mechanism, the paper describes the flow phenomenon sequences during the buzz cycle by means of the shadowgraph pictures and via high frequency pressure transducers. Results showed that the pressure inside the intake decreases and increases sequently during the buzz cycle. The intake becomes almost empty (its mass flow rate decreases) as the shock wave moves upstream toward the intake tip. When the shock waves stand at its most upstream location, the pressure inside the intake reaches its minimum value. This low pressure condition causes the shock wave to move toward the intake and consequently the intake pressure increases again. As the pressure inside the intake increases, the shock wave moves upstream. The intake pressure reaches its maximum value when the shock wave stands at the intake entrance and the buzz cycle is then completed.