نشریه علوم و فناوری فضایی
سال سوم شماره 1 (پیاپی 6، بهار و تابستان 1389)

In this study, the combustion process of gaseous Oxygen on the surface of HTPB solid fuel has been investigated. To simulate the chemically reactive flow, Navier-Stokes equations and species transport equations were solved using LU-SW implicit scheme. Modeling this kind of combustion process demands a deep understanding of the pyrolysis phenomenon on the solid fuel surface. Experimental studies conducted in this field show that the main gaseous product of the pyrolysis process is C4H6. An experimental equation which is dependent to the temperature of the fuel surface is used to determine the gas production rate during pyrolysis process. The temperature of the fuel surface can be obtained by applying energy equation in gas-solid interface. The combustion process of gaseous Oxygen and C4H6 has been described by two quasi-global chemical kinetics models. According to the obtained results, the main characteristic parameters of combustive flow such as the flame temperature and mass fraction of chemical species are strongly affiliated to the applied chemical kinetic model. Finally, the results of modeling based on two different models of chemical kinetics are presented and solid fuel surface regression rate is compared with other numerical results.

Separation system is one of the main sub-systems in every space device and missile. This system is used to separate active and inactive sections in a missile. The separation process should be done accurately. A separation system explained in this paper was designed and produced based on mission, rocket configuration and manufacturability. This system uses explosive bolts and spring mechanism. This separates accurately payload from motor in a sounding rocket. Aerodynamic forces were used to determine structure strength. This structure designed based on at least weight and drag force. Several tests were done to evaluate this system performance. The tests results confirm this system capability.

Designing a manned spacecraft carrying one or two persons to low Earth orbits needs to recognize system level requirements and acquire technical data developed in this eria. Revising manned spacecrafts’ characteristics leads to recognize system level requirements and achieve applicable results which are needed to design and development of such a spacecraft. Manned spacecraft characteristics comparing charts and figures show a roughly analogous pattern in terms of mass and dimensions and confirm the parallel subsystems have similar performance.

In this paper a new guidance law is proposed to guarantee the stability of the guidance loop considering first order pursuit dynamics using short time stability theorem. As homing guidance is operates over a finite time, short time stability criterion which is defined over a specified time interval can be used effectively in guidance loop stability analysis. Proposed guidance law utilizes line of sight angular rate and pursuit acceleration measurements. Stability region which depends on the pursuit dynamics and guidance gains is an analytical expression in terms of time to go. Stability condition of the new guidance law is less conservatism than classical proportional navigation guidance law.

In this paper, the complete block diagram of the imaging payload of a spin satellite capable of real time imaging is designed. Because of the satellite spin, the system needs to recognize the suitable camera angle in order to start imaging. The angle is the starting point of the observation of the part of the earth to be imaged. In this paper, at first the suitable imaging method and detector for this kind of satellite are elected and then the angle and the time of the spin camera imaging and the necessary number of lines and pixels are calculated. If the system is also capable of real time imaging, the captured images should be transmitted to the earth station before the next imaging starts. The completion of the above scenario needs a complete and parallel relationship between the satellite image payload subsystem and other subsystems such as power, communication and specially satellite on-board computer. For imaging and transmission, image payload status information such as temperature, voltage and current should be sampled and transmitted to the on-board computer for processing. Also this information should be attached to the image frames and transmitted to the earth station. All this processing is summarized into time pulses with exact timing between subsystems. Because of resource limitation in a space mission, satellite systems design must have the minimum mass, power and cost. But these shouldn’t cause the efficiency and specially system processing speed to decline. Imaging payload with real time capability needs a high processing speed requiring high resource utilization. In this paper, an imaging system is designed with the mentioned characteristics based on FPGA high parallel processing speed but having low mass, volume and power.

In the present study, in order to choose the suitable heat shields a comprehensive investigation was performed. Therefore, different properties of each type of heat shields including ablative and thermo-physical properties were measured separately. Finally, the obtained properties were compared. The studied heat shields are phenolic resin composite reinforced by ceramic, asbestos AAA and C cloths. The experimental investigations consist of oxyacetylene standard flame test, specific heat capacity, thermal analysis and thermal conductivity. In addition, in order to compare the thermal efficiency of mentioned heat shields, the temperature on the back surface of each sample subjected to constant heat flux was measured. The results showed that the specimen with 8 mm thickness reinforced with ceramic cloth, which was subjected to 2500 kW.m-2, reduced the temperature from 1550 to 100oC. Therefore, the phenolic composite containing ceramic cloth is the best option.

Orbital transfer has a significant role in any space mission. This transfers generally categorized in impulsive and continuous maneuvers. An important challenge is fuel consumption in the maneuver. This problem is considered as a required 􀎿􀝒 problem. Hence, minimization of 􀀃􀎿􀝒􀀃 means minimization of fuel consumption orbital transfer. In simple cases, the problem has closed form solution for example transfer between coplanar circular orbits or transfer between coplanar coaxial elliptical orbits. The conventional methods cannot solve complex cases of initial and target orbits. In this paper the impulsive optimal transfer between two coplanar- noncoaxial elliptical orbits is considered. The numerical solution of optimality nonlinear equations is necessary to obtain the solutions of complex problems. According to nonlinearity of equations two issues arise, firstly numerical solution is sensitive to initial guess, secondly the local minimum solutions only may be find. In this paper some equations have been derived that using them behavior of required 􀀃􀎿􀝒 function can be investigate based on various values of independent variables and can be find the boundary of global solution. In this way one can be determined a reasonable and proper initial guess for nonlinear solver. The proposed methodology is applied to an example and the results are provided. The results include the local and global solutions and they show a good ability of the proposed method.

Development of halo orbits and their associated invariant manifolds are investigated. Halo orbits play a fundamental role in complex space mission designs. In essence, halo orbits are periodic solutions of the restricted three body problem (R3BP) determined under specific initial conditions. In this paper, the symmetric property of the nonlinear R3BP governing differential equations is utilized in order to obtain the desired initial conditions. In this regard the differential correction technique and the state transition matrix are used to generate the halo orbits. The differential correction technique, based on the Newton method, is an effective tool for solving two point boundary value problems. In addition to generate the stable and unstable manifolds, the initial conditions are perturbed in the direction of Eigenvectors and the equations of motion are integrated for an arbitrary time interval.