Feasibility study of a composite structure for a LEO satellite

Composite materials have been studied and used for decades in the space industry because of their specific properties, which allow the designing of satellite structures that are both lightweight and robust. The development of the Low Earth Orbit (LEO) small satellite market over the last two decades has led to the need to design new structures in order to address the challenges of miniaturization and maximization of payload capabilities. This thesis has studied whether composite structures are suitable to endure the low Earth orbit environment and whether they can offer substantial benefits compared to traditional aluminum structures. Among composite materials, the focus has been put on Carbon Fiber Reinforced Polymers (CFRP) because of their high performance and versatility. It has been underlined, through several academic sources and technical reports, how CFRP structures can withstand LEO conditions, addressing concerns related to atomic oxygen exposure and hypervelocity impacts. Academic literature highlights CFRP sandwich panels for their superior tolerance to impacts over monolithic panels and for potential multilayer shielding concepts. Additionally, CFRP demonstrates potential advantages over aluminum for radiation shielding, crucial for long-duration missions and future space exploration efforts. Hawk for Earth Observation, a microsatellite designed by Argotec for the Iride constellation, has been the subject of a feasibility study for a composite architecture. An analysis of typical configurations for a small Earth observation satellite was carried out, along with a study of composite manufacturing technologies. The most fitting solution for the design of the satellite structure is a central-core CFRP monocoque cylinder and CFRP sandwich panels as external enclosures. The dimension of laminates and the number of needed inserts have been identified as crucial aspects for determining the convenience of a composite design. It has been found the minimum panel size under which the benefit of sandwich architecture becomes negligible, considering the number of inserts estimated to be necessary. From this analysis, it has been observed that the studied satellite falls into the design range where a material change would provide substantial weight reduction and increased stiffness. Structural analyses have been carried out to verify the efficacy of the new structure. The resulting displacements and stresses were not critical for the satellite, reaching a minimum factor of safety of 7,08 for the aluminum honeycomb. The first natural frequency has shown a 19,7% increase compared to the existing structure. On the other hand, the new positioning of subsystems on the external panels led to potentially hazardous acceleration for the electronic components when exited with the launcher random vibration spectrum. The proposed solution included using dampers for critical components while highlighting the need for further studies on an optimal internal configuration. Finally, the presented design has observed a weight reduction estimated to be around 37,6% concerning the structural mass and a 17,2% decrease concerning the total wet mass.