This research discusses the effect of simultaneous usage of PROPELLANT Utilization (PU) system and Flight Apparent Velocity Regulation (AVR) system. These systems were used for sending OBC commands to engine for adapting the engine working regime with flight conditions and fame to active control systems. Each of the PU and AVR systems has an effective role in access to final parameters such as mass and velocity at the end of active phase and simultaneous usage of these systems leads to increased range accuracy and payload mass. We study these effects on final parameters in this paper. Therefore, with dynamic simulation of liquid PROPELLANT engine during active phase in flight simulator, sending commands of these systems to change the engine working regime are provided. For a specific mission, results show that using the PU, range increased and presence of AVR is assisted to reach this range in front of disturbance during the flight. Another important result of this research is the payload mass increased for a specific mission with simultaneous usage of PU and AVR systems.

A theoretical study analyzing three-dimensional combustion acoustic instabilities in a liquid PROPELLANT rocket engine combustor has been conducted. A linear theory based on Crocco's pressure sensitive time lag model is used. To apply this theory, the combustor is divided into two main components, including the combustion chamber and the converging part of the nozzle. The assumption of concentrated combustion zone is used and the governing perturbation equations describing oscillations of flow variables are considered. To solve these equations appropriate boundary conditions at both ends of the combustion chamber are req1lired. Combustion zone boundary condition at one end and the nozzle admittance relation at other end are used. To obtain the nozzle admittance the three dimensional flow perturbation equations are solved in the converging part of the nozzle. This approach is capable of predicting acoustic stability behavior of a combustor at a wide range of Mach numbers and frequencies. Also, this analysis enables the rocket engine designer to observe the effects of different parameters such as nozzle entrance Mach number, chamber geometry, nozzle geometry. and gas properties on stability characteristics of an engine combustor. In case of instability observation; one can predict the acoustic mode which causes the instability and achieve an optimum design before conducting any expensive and time consuming experimental tests. This paper presents the stability analysis results and a parametric study of the effect of design parameters on stability characteristics of a typical combustor.

The effects of liquid PROPELLANT rocket engines thrust termination transients are important in achieving the launch vehicle’s desired final velocity with the required precision. In this paper, a mathematical model has been developed to predict the changes in the combustion chamber pressure and the related cut-off impulse based on the physical aspects of engine’s components. The modeling is divided into four steps: (1) from issuing the cut-off command to the activation of valve (2) from the time of executing the stop command until the end of the operation of the cut-off valves, (3) from the time that after the valve is shutt closed until the combustion chamber is extinguished and (4) simulation of the phase change and PROPELLANT components evaporation in corrugates. Results suggest that the duration of the first two steps have a significant effect on increasing or decreasing the amount of the thrust force and the 4th step’s thrust is less than 10 percent of nominal value while, the most thrust fluctuations appear in this step.

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.

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.

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.

Gel PROPELLANTs, in comparison with liquid PROPELLANTs, have some significant advantages. Atomization and then combustion of gel PROPELLANTs can provide required thrust for propulsion of missiles. Characteristics of atomization and combustion of PROPELLANT depends on its rheological properties. In this paper the rheological properties, atomization pattern and combustion processes of liquid and gel PROPELLANT are presented. Gel PROPELLANT is a shear thinning fluid, so its viscosity is decreased with increasing shear rate. Thus, suitable injectors for this kind of PROPELLANT are those, that are able to apply high shear rates on PROPELLANT during passing through the injector. Smaller diameter of atomized droplets, lead to faster evaporation and more efficient combustion, so the combustion process can be completed in a shorter length of thrust chamber.

In this paper, a fault in the main duct of oxidizer in liquid rocket engine has been simulated. The procedure is based on a nonlinear mathematical model that dynamically describes sound operation of engine. In other word, when engine operates abnormally during a fault, fault descriptive equations are added to the engine equations. Differential equation system involves linear and nonlinear equations that describe sound and fault operation of engine. This system consists of 34 algebraic and 59 differential equations. The fourth order Runge-Kutta method has been used for numerical solution of equations set. Finally, results of engine hot test are compared with model and shows acceptable accuracy of simulator code.