Structural analysis and design of composite 3U cubesats

In this thesis study, structural analyses of nano satellites which have different materials and design configurations, were made under loads that occur in the time from production to end of the mission life of satellites, and the obtained results were compared. The nano satellites were designed with the same dimensions. With the thesis study, determination of the structural behavior of the satellites, which were designed by using different materials, against the loads, is aimed. First of all, the dimensional limitations of nano satellites, which are examined in the thesis, were mentioned. Properties of the production, transportation, launching and the space environment conditions, and these environments effects on the structural subsystems of the satellites were explained. Launch process contains several basic stages. These are liftoff and high dynamic pressure effects phase, stage jettison, payload fairing jettison, stage ignition and satellite separation. During the launch stages, high mechanical loads like acoustic, shock, random vibration and quasi-static loads occur. Satellites should withstand these loads without any structural failure. Also after the launch, satellites work in the space environment, which is not very conducive to structural materials, especially carbon composites. This environment becomes very harmful to materials with some properties like electromagnetic radiation, high temperature difference, atomic oxygen and vacuum. Particularly, temperature difference causes thermal loading on satellite components. In the thesis, these loads, which affect the satellite structures before and during the mission, were identified. Subsequently, CAD drawings of nano satellites were done by using CATIA V5R21 software. At design stage of satellites, the "CubeSat" standards were taken into consideration. When defining the dimensions these standards were used. The basic structural properties and other important specifications of "CubeSat" are determined by California Polytechnic State University (Cal Poly) and Stanford University. CubeSat concept has same dimensions in all sides of satellites as 100mm. In other words, CubeSats are cube shaped and generally used with these dimensions. However, double and triple configurations are possible. In this thesis, triple configuration satellites, which are called as 3U CubeSats, were preferred. At the end of the structural strength analyses, in order to achieve realistic results, nano satellites should be modeled with all subsystems including structural subsystems. During modeling, the component masses of these subsystems are gaining importance. Because of this reason, a task definition was determined for nano satellites and sample components were selected for essential subsystems. One of the critical stages of satellite design is selection of the materials. 3U CubeSats, which were examined in thesis study, have 340.5mm side length in z direction, because of this reason more bending moments can be occur compare to 1U CubeSats. Thus, selection of the materials gained more importance. The aluminum and carbon composite materials, which are used in main structure of satellites, were determined between the options. In general, aluminum 7075, 6061 or 5052 are used for the main structure. In the thesis, Al 5052-H32 was chosen as main material for the aluminum satellite. Also, in order to determine carbon fiber and sandwich core materials a literature research was made. As a result of this search, the unidirectional prepreg USN 150 and wowen prepreg T650-35 3k/976 were selected for monolithic parts and face sheets of sandwich panels. In addition, CR-III Aluminum 5056 honeycomb was chosen as core material of sandwich panels. Then developed protective processes against the harmful effects on the materials of the launching and space environment conditions were explained. Especially, carbon composites are affected space environment conditions negatively. Carbon composites react with atomic oxygen while aluminum panels are nearly unaffected. Also high temperature differences can cause micro cracks in carbon composite resin. In order to prevent these effects use of protective coating was considered. In thesis, aluminum foils were preferred as protective coatings on carbon composite panels. By using this method, long term protection was aimed on the structures of satellites for all environmental conditions. After the material selection, shape of the top, bottom and side panels were decided. It is very important because these panels constitute the basic structure of satellites. In the thesis, two different panel forms were designed. The first type panels have big windows on their surfaces. These panel designs were used in aluminum and one of the monolithic composite CubeSats. Other panels do not contain any window on their surfaces and were utilized in the other monolithic composite and sandwich composite CubeSats. Also, the second support structure was designed for all CubeSats. This structure includes four stainless steel endless screws and aluminum spacers. The endless screws were used in order to fasten the printed circuit boards to each other and to the CubeSat panels. Spacers prevent the motion of subsystem components during the launch and on the orbit. In addition, layer properties of composite parts are very important. These properties can be written as, type of the lamina that used as a layer, thickness of one layer, fiber direction of a lamina and layer numbers. All properties affect the strength of composite structure. In this thesis, unidirectional and wowen carbon fiber/epoxy composite perepregs were used as a lamina in monolithic parts and face sheets of sandwich panels. Also, wowen carbon fiber composites can be called as bidirectional because of their fiber arrangement. In addition, aluminum honeycomb was selected as the core material of sandwich parts. Two satellites were designed using with only monolithic panels. The other composite satellite includes both monolithic and sandwich parts. In order to achieve maximum strength in the satellite structure, locations of the unidirectional and wowen prepregs were determined carefully. Unidirectioanal carbon fiber/epoxy composites were located on the side panels and wowen composites were used on the top and bottom panels. This usage was preferred in all composite structures. After the structural design stage, determination of the structural loads acting on satellites during launch is required. Therefore, the methods, which are used for calculation of shock, quasi-static and random vibration loads were explained. Calculation method of the resultant loads, which are applied during the static analyses, was told. After load calculations, the classical lamination theory and the post-processing methods, which were applied on carbon composite structures, were explained. These methods were used in order to determine the failure modes, which can occur on monolithic or sandwich composite parts. The CAD drawings of nano satellites, which are examined in the thesis, were transferred to the ANSA meshing software. By using this program all CubeSat structures including panels, endless screws, spacers and printed circuit boards were meshed separately. Then all of them were extracted to MSC Software Patran 2010 finite elements program. The required analyses were done with this software. According to materials, which are used in their structures, the material properties were identified into Patran. The structural analyses were done firstly for launch conditions and then for space environment conditions on the satellites. The MSC Software Nastran 2010 program was used in order to analyze the satellite structures. In either case, primarily boundary conditions and loads were defined. Afterwards, modal analyses were performed in order to calculate the stiffness of satellites, which are located into P-POD during the launch. At the end of the modal analyses, satellites natural frequency values and the mode shapes at these frequencies were obtained. The calculated natural frequency values of satellites are provided the required minimum stiffness value of launch vehicle. The random vibration loads which occur in launch conditions were transformed to g-loads by using Miles method. These g-loads were calculated with the natural frequency values of satellites at different modes. Then the transformed random vibration loads were combined with the quasi-static loads, which are obtained from launch vehicle user guide. The combined loads were applied on the finite element models of satellites. After all boundary conditions and loads were defined on the models, static analyses were done for all satellites and obtained results of aluminum satellite were presented. In order to evaluate the failure situation of carbon composite structured satellites, some post-processing methods were applied on carbon composite components. At launch conditions, according to these post-processing methods, any structural failure did not occur on satellites. The similar procedure was followed in analyses, which were done for space environment conditions. However, different boundary conditions and loads were identified in finite element models of satellites from launch conditions. In these analyses, satellites were not bonded anywhere and they were modeled as free like they are moving in their orbit. Also in space environment there are not any mechanical loads caused by launching, for this reason only thermal loads were applied on satellites during static analyses. At the end of the analyses, structural failure occur on the satellite, which made of aluminum. However, the carbon composite structured satellites were post-processed and unlike aluminum one any structural failure did not appear on their components. As a consequence, with this thesis study it was shown that the carbon composite structured satellites, which were designed in "CubeSat" standards, could work under mechanical and thermal loads without any structural failure. All results of static and dynamic analyses were seen in expected levels. Also with the new design of carbon composite satellites, mass reduction was occurred in structural subsystem. Carbon composite structured satellite, which has sandwich parts, 26% lighter the aluminum one. In addition, the monolithic composite satellite, which has panels without any window, was extenuated 27% by using laminated carbon composites in all primary structure. Besides the other monolithic composite satellite, whose panels have windows, 6% lighter than the aluminum CubeSat. The windowed composite satellite is heavier than other composite ones because external solar panels were used on it. However, windowless CubeSats do not require solar panels. Solar cells can be mounted on CubeSat panels.