نشریه علوم و فناوری فضایی
سال شانزدهم شماره 1 (پیاپی 55، بهار 1402)

In this paper, a step-by-step laboratory procedure for performing a satellite's payload’s alignment measurement is presented. Four highly accurate theodolites are used along with two or more alignment corner cube to accurately extract the final attitude. Theodolites are arranged around the satellite in such a way that they have a clear direct view of the alignment cubes mounted on the payload and the satellite. Two theodolites should point to the payload’s alignment cube and the other two theodolites must point to the satellite’s alignment cube. Each theodolite must see at least one other theodolite, directly. Finally, by forming the coordinates systems of the payload and satellite in the theodolites coordinate system along with using the coordinate transfer matrices, the payload alignment correction matrix will be extracted in detail. The total method accuracy is within the order of few arcseconds.

One of the passive components of the satellite Thermal control subsystem is multilayer insulation. In order to prevent air from being trapped between the multilayer insulation layers, which causes the thin layers to inflate and disintegrate during satellite launches, holes are made in the layers. These holes in different layers may not be aligned due to heat transfer problems as well as manufacturing constraints. For maximum thermal efficiency of thermal insulation, gas outlets must be designed to have the least resistance to exhaust gas flow, because the air trapped between the layers will greatly reduce the insulation efficiency by leaving a convective heat transfer path between them. In this article, different perforation matrix that have been used in articles are reviewed. By analyzing the computational fluid dynamics of gas outflow from these insulators, the effect of various parameters has been studied.

The main purpose is to introduce the performance system design and optimization method of aerospike nozzle for different aero-space conditions. For this purpose, some of the important parameters of the aerospike nozzle structure and cold flow condition tests in the nozzle optimization are studied. The methods of designing the Aerospike nozzle and its governing equations are described and the proposed design model is described and important factors are expressed in this type of nozzle. therefore, the design of a complete nozzle is made by aerospike and is supported by an existing design sample. Then, in order to optimize the nozzle, three cuts of 20%, 40% and 60% of the nozzle end are analyzed. The standard for comparison and optimization in these three slices is the Mach number of the output current. The results of this comparison show that the most efficient aerospike nozzle is a 40% cut nozzle based on the flow charts and contours of this aerospace nozzle.

Monopropellant thruster are used to inject a satellite into orbit or control its position on three axes in space missions. One of them is hydrazine thruster which is widely used. In this research, design of the injector, decomposition chamber and nozzle of a 10N hydrazine monopropellant thruster have been performed. The capillary injector was designed using Fluent software for this thruster which was able to supply the mass flow rate of the thruster (5 gr/sec). The decomposition chamber contains catalyst granules and its dimensions were selected based on the complete decomposition of hydrazine. The nozzle was designed by RPA software. The validation of the design with RPA software was checked by a numeric code. This code was able to calculate the dimensions of the decomposition chamber based on the amount of hydrazine decomposition. Accordingly, the results of both design methods are strongly consistent with each other. At the end of the design, the final thruster design and drawings were prepared.

The aim of the present research is to obtain the ability to use the cryogenic propellant engines on a laboratory scale. In this regard, it is necessary to build some experimental motors and investigate the their performance parameters. The liquid oxygen as a common oxidizer and ethanol as a green fuel have been selected as propellant components. The engine is designed to produce 400 kgf force at the nominal condition. The pintle type injector has been chosen in which liquid oxygen and fuel are flowed in the axial and radial directions, respectively. The combustion chamber has been protected against overheating by applying the regenerative cooling. However, the laboratory feature of the engine design has provided the using of water instead the cooling propellant. All main components of the engine such as injector, igniter, and flow controllers, are examined by the cold tests. A comprehensive test facility is designed and set up for hot fire tests in which the performance of almost all parameters can be evaluated. Fifteen fire tests have been performed. Maximum obtained pressure and evaluated combustion efficiency were about 75% of design values.

With the advent of GaN technology, achieving microwave power with high efficiency by solid-state devices has become more available. Therefore, the use of SSPA amplifiers with GaN technology in satellites, especially LEO satellites, has been considered. space radiation can affect the performance and reliability of components in space systems, which needs to be investigated. One of the most important technologies that can be affected by radiation effects is GaN transistors. In this paper, the effect of TID on GaN transistors in the SSPA amplifier board is investigated. Since commercial components have been used in the engineering sample of the SSPA amplifier and the calculations obtained from the RDM estimates under the worst conditions show that it is necessary to conduct a test for these components, the radiation resistance test was performed for this amplifier. The results of the test conducted in this article show that the SSPA GaN board has radiation tolerance up to a dose of approximately 16 krad. Therefore, mismatched GaN transistors are resistant up to this amount of dose. This is while the sequencer board actually has less tolerance than 5.5 krad

To test the Attitude Determination and Control Subsystem of a satellite, it is necessary to have an attitude dynamics simulator, and the simulator must be in a balance condition. Disturbances on the balance system in the simulation include deviations caused by the difference between the center of mass and rotation, as well as the movement of two horizontal actuators. The movement of two horizontal actuators is a factor for rotational and vortex motion. In the simulation of experimental models, PID control coefficients are also used to control three axes. The balance system actuators include moving masses and reaction wheel that are installed around the horizontal and vertical axes, respectively. To validate the results, a hardware sample has been developed for laboratory tests. Using the sampling time, models and experimental coefficients, the hardware reaches the accuracy of 0.2 and 0.5 degrees in 25 seconds, respectively, which indicates a suitable accuracy for balancing the simulator of the CubeSat attitude.