The pulsed plasma Thruster has been the first thruster in the space mission. In the thrusters, based on discharging electrical capacitor and passing high current between the anode and cathode, fuel is isolated, then by using self-field magnetic and applying Lorentz force to the plasma particles and with accelerating them the thrust force is generated. In the research a one dimensional Pulsed Plasma Thruster has been investigated. The applied numerical solution is based on Einfeldt, Harten, Lax, Van Leer, (HLLE) numerical method which has the adequate accuracy. In the simulation Hall effects, ionization process, heat transfer, and viscosity have been neglected. Governing equation for a magnetic accelerator has been solved. The represented solution results include distribution of density, velocity, pressure, and magnetic field during the magnetic accelerator which compared with similar numerical results was satisfactory. A Pulse Plasma Thruster has been numerically analyzed. The results related to density, pressure, magnetic field, and velocity curves have been compared with the desire physical behavior, which were satisfactory. Also graph of distribution of Teflon temperature after reaching to the ablation temperature, by applying heat energy from plasma area, has adequate compatibility with reference.

A systematic design of the Solid propellant rocket; Analysis of characterizes and components; Manufacture of a porotype and flight test are the main purposes of study. In rocket motor design process, the zero and one dimensional interior ballistic flow solution are used. The numerical solution is based on iterative approach for propellant combustion conditions and pressure estimation. propellant grain geometry, combustion chamber, nozzle, motor bulkhead, joints analysis and sealing are most important subsystems; so the challenge is to develop a model for their convergence. The other highlight of this approach are parts building; make Solid propellant and finally assemble the motor. This motor was used for real flight test of the sounding rocket. The experimental data and scientific process have good convergence and its results usable for other engineering research.

In this study, the simulation of internal ballistics and the exhaust flow of a Solid propellant engine have been investigated. The flow field considered in this problem including the cylindrical grain and the Solid wall of the engine and the flow at the nozzle outlet have been studied. The governing equations are used in an axisymmetric and compressible form to model the fuel combustion in non-erosive combustion mode. In the present study, cylindrical type of grain fuel is considered, which is one of the most useful types of fuel grains in terms of simplicity in production. In order to obtain the results from the simulation model which are close to the real data, used UDF modules in the Fluent software to describe some specific quantities and phenomena of the problem. Finally, the results obtained from the simulation in different modes of fuel burning are examined and show the accuracy of the numerical model by entering the parameters affecting the fuel burning.

In this paper, specific grains burn back is presented by new geometrical method. The software is developed for Wagon wheel grains and 3D grains. Rapid Solid motor ballistic simulation code produces required charts with considered nozzle geometry. Presented method used geometrical introducer point (GIP) to produce the various grains.Limitations and configurations are simply modeled. Thrust, total pressure and temperature are illustrated as web burned changing. The results confirm the performance of the developed algorithm for mentioned grain analysis. Lower time processing and rapid ballistic analysis are the benefits of the presented algorithm. Finally results of the burn back analysis code and the internal ballistic simulation code are evaluated by some other existent codes and real cylindrical grain test.

The effects of erosive burning of Solid propellant in a prototype motor have been investigated through "Interrupt Burning" method. The motor is composed of three internal-external combustion cylindrical propellant in motor section with the L/D ratio of 5 and one two-ended internal combustion cylindrical propellant in test section. Erosive burning will be examined taking into consideration of three working pressures and mass flux that occurred in the port. In this paper, a new equation will be introduced in which the effects of mass flux, burning rate, combustion chamber pressure and motor dimensions are included. The prototype motor will be tested and then the obtained results will be used to determin the parameters of the proposed. Therefore an applicable equation in grain designing will be obtained.

In this paper a new method for guidance of Solid propellant and aerodynamic control missile is introduced. In proposed method after achieving maximum acceleration، flight path angle of missile changes in order to impact to target. In this method flight information such as time history of magnitude of velocity and position vectors are saved in the table before firing. For calculating difference of nominal and disturbed parameters new concepts are used. Corrected values of nominal parameters are used in guidance after firing to guaranty impact to target. The flight simulation of proposed method is compared with functional guidance method and the effect of the proposed method is shown. Design verification is done in base of simulation in presence of flight disturbances.

There are several sources for pressure ascillations in Solid propellant rocket motors. Oscillatory flow is one of them. Free shear layers in motor flow field cause vortex shedding. End edges of propellant grains baffle edge in two-segmented motors are samples of such zones. These vortices move from their forming points and strike the field walls. The kinetic energy of vortices change to pressure, forming acoustical pressure ascillations. Acoustical characteristics of pressure ascillations such as frequency and amplitude change with the gradual change in the internal geometry of the motor. In this paper, the interaction between mean flow and acoustic field in a Solid propellant rocket motor is studied numerically. Roe's flux function in an unstructured grid strategy for solving compressible viscous flow equations shows large changes in frequency of pressure oscillations in motor. Six different motor geometries are used for simulation of motor internal geometry at different burning times and grain configurations. Using this methodology, the frequency and intensity of pressure waves are well predicted. It is also shown that frequency jump from second longitudinal mode to the first is formed as a result of changes in the internal geometry.

Internal ballistics of Solid rocket motor is simulated by solving the axisymmetric Navier-Stokes equations on a moving unstructured grid. Then, the flow velocity, pressure, temperature, and other variables are computed for problems with complex geometries and boundary conditions. The dynamic (moving) mesh is used to model flows where the shape of the domain changes with time due to motion of the domain boundaries (burning surfaces).Updating of the mesh is handled automatically at each time step, based on the new position of the boundaries. This study showed that CFD with unstructured dynamic meshing can be a powerful tool for simulating internal ballistics of Solid rocket motors.