This paper presents the theory and design of a semi-empirical atmospheric density model based on data from the MSIS-86 model. Created as part of ongoing research into optimal guidance laws for nanosatellite applications, this model focuses on being computationally lightweight, while providing reasonably accurate atmospheric density predictions at geometric altitudes ranging from 0-1000 km. The model is validated against data from existing analytical and empirical atmospheric models. It is then implemented in a variety of orbit and attitude propagation environments in Matlab-Simulink to assess its stability, validity, and computational footprint. The orbital elements from each simulation were compared against those obtained from baseline simulations run using the Naval Research Lab (NRL) MSISE-00 model. The results show good agreement with the baseline simulations, while indicating a significant reduction in computational run time.

This paper presents the theory and design of a torque-optimal guidance algorithm for CubeSat applications. CubeSats and nano-satellites provide mission-flexible low-cost platforms for the academic and scientific communities to conduct cutting-edge research in the harsh environment of space. The mission life of nano-satellites may be limited by the attitude actuators, and it is therefore beneficial to reduce torque and angular momentum usage during reorientation maneuvers. The algorithm focuses on being computationally lightweight and robust, while including the effects of gyroscopic moments, environmental torques, and degraded reaction wheels. Results indicate that this torque-optimal guidance algorithm demonstrates substantial improvements in performance and pointing accuracy over an Eigenaxis controller for similar maneuvers, with low to moderate computational overhead. In doing so, it presents a significant advancement towards the development of intelligent GN&C systems for small satellites.

CubeSats and nano-satellites provide flexible low-cost platforms for the academic and scientific communities to conduct cutting-edge research in the harsh environment of space. The mission life of nano-satellites is often limited by the attitude actuators, and it is therefore beneficial to reduce torque and angular momentum usage during reorientation maneuvers.In this capacity, a computationally lightweight torque-optimal guidance algorithm was formulated, solved using pseudospectral methods, and validated in a MATLAB-Simulink environment. A low-computation atmospheric density model, developed in support of this research, was extensively validated via performance assessment of passive CubeSat aerostabilization.Results indicate that this torque-optimal guidance algorithm demonstrates substantial improvements in performance and pointing accuracy over an Eigenaxis controller for similar maneuvers, with low to moderate computational overhead. In doing so, it presents a significant advancement towards the development of intelligent GN&C systems for small satellites.iii I n m em o r yo fC o l um b i aa n dC h a l l e n g e r . T h i s r e s e a r c h i sd e d i c a t e d t o t h o s ew h oh a v e t a u g h t m e t h a t t r u eh a p p i n e s s i n l i f e i s n o ta b o u tr e a c h i n gad e s t i n a t i o n ,b u t i n m a k i n g m em o r i e s f r omt h ea d v e n t u r e sa n d c h a l l e n g e sd u r i n gt h e j o u r n e yt o w a r d s i t . S i c i t u ra da s t r a i v

This paper presents an extension to the SPeAD-M86 model by Kedare & Ulrich (2015) by incorporating variations in atmospheric density resulting from the 11-year solar cycle, as quantified using the F10.7 index. It focuses on utilizing sinusoidal and exponential piece-wise functions to estimate temporal density changes in the atmosphere at geometric altitudes ranging from 0-1000 km. The model is validated against data from existing analytical and empirical atmospheric models. It is then implemented in a Matlab-Simulink orbit and attitude propagation environment to assess its stability, validity, and computational footprint at various instances in the solar cycle. The orbital elements from each simulation were compared against those obtained from baseline "truth" simulations run using the Naval Research Lab (NRL) MSISE-00 model. Results indicate improvements in accuracy compared to the SPeAD-M86 model with minimal increase in computational run time.