A cube satellite (CubeSat) is a pico-class miniaturized satellite with a dimension of 10 × 10 × 10 cm3 and a mass less than 1.33 kg with respect to a standard size of one unit (1U). In addition, the CubeSat can be expanded to more than 2, 3, 6, and 12U, which is the basic size depending on the mission purpose. CubeSat is vert flexible to exploration space mission and on-orbit scientific verification that assists the next generation of scientist and engineering to complete all phases of space mission. Recently, Due to the miniaturization of electronics and the increase system integration density, the CubeSat had been used for increasingly complex missions. Also, High-performance equipment such as cameras and sensors will be mounted on satellite bodies according to the mission of the CubeSat, which has resulted in the overall increase in the power consumption of the satellite system. However, it is difficult to meet the power consumption by attaching the solar cell on the surface of the CubeSat. To enhance the electrical capability of a CubeSat, Deployable solar panels are required, which would increase the surface area of the CubeSat for the installation of solar cells and would help orient the panels to face perpendicular to the sun for effective power generation.
Recently, various holding and release mechanisms has been developed for separation of the deployable solar panel. In general, Pyrotechnic devices are widely used to commercial satellites. The explosive restraint system has high frequency separation reliability based on its high tightening power and performance in the application of accumulated space missions, there is a problem that does not satisfy the design requirements of the CubeSat due to the high shock level involved in the operation of the restraint separation device. To overcome the drawback of pyrotechnic devices, non-explosive devices based on Shape Memory Alloy are being applied to the actual mission. However, shape memory alloy device has the advantage of low shock level compared to Pyro, there are difficulties in satisfying limited size, including development cost, for applying cube satellite with expensive equipment. Currently, various holding and release mechanisms are used in many areas of the space industry, but there is limitation in the application of the CubeSat, because of the shock level and high cost. Therefore development of holding and release mechanism for CubeSat is steadily required to replace disadvantages of pyro and shape memory alloys. A nichrome burn wire cutting method is extensively used for the release of mechanical constraints on the deployable solar panel of the CubeSat owing to its low cost and simplicity. The mechanical constraint was released by cutting the nylon wire triggered by a resistor.
In this study, the nichrome burn wire triggered Holding and Release Mechanism using a spring-loaded pogo pin is proposed and this is developed for the separation device of solar panels. A pogo-pin is used in electronics to establish a temporary connection between two printed circuit boards. This is possible to the application of separation devices because it functions as an electrical interface to provide power, separation spring to initiate reaction force and the status switch to determine a deployment status of the solar panel. The advantages of the mechanism proposed in the study include increased loading capability in the in-plane and out-of-plane directions of solar panels, synchronous release of multiple panels, low-shock levels, and handling simplicity during the tightening process of the nylon wire. The release function tests of the mechanism was verified by performing a separation test under various conditions. These test results demonstrated that the proposed mechanism verified the feasibility of using pogo-pin as a holding and release mechanism of solar panels for CubeSat applications in accordance with the proposed design. Also, pogo-pin based Holding and Release Mechanism was verified through the qualification level of the sinusoidal and random vibration tests. The Thermal Vacuum test was also performed to determine the survivability and reliability of the Holding and Release Mechanism in the simulated on-orbit thermal environment. The test results demonstrated that the proposed solar panel Holding and Release Mechanism was certified for use in actual space missions. The simple, reliable, and inexpensive Holding and Release Mechanism often contributes to significant cost reductions to a CubeSat program and aids in the success of space missions.