Though the Aerospace industry is moving towards small satellites and smaller sensor technologies, sensors used for close-proximity operations are generally cost (and often size and power) prohibitive for University-class satellites. Given the need for low-cost, low-mass solutions for close-proximity relative navigation sensors, this research analyzed the expected errors due to near-field effects using a coarse sun sensor as part of a 6-degree-of-freedom (6-dof) solution. To characterize these near-field effects, a test bed (Characterization Test Apparatus or CTA) was proposed, its design presented, and the design stage uncertainty analysis of the CTA performed. A candidate coarse sun sensor (NorthStarTM) was chosen for testing, and a mathematical model of the sensor’s functionality was derived. Using a Gaussian Least Squares Differential Correction (GLSDC) algorithm, the model parameters were estimated and a comparison between simulated NorthStarTM measurements and model estimates was performed. Results indicate the CTA is capable of resolving the near-field errors. Additionally, this research found no apparent show stoppers for using coarse sun sensors for 6-dof solutions.

This thesis describes technologies developed for nanosatellites at the Space Flight Laboratory. A critical ground station component, the Terminal Node Controller, was upgraded in order to support Generic Nanosatellite Bus and future missions. Sun sensor requirements and operation were reviewed, followed by details of the author's work in executing the acceptance testing on these parts, including thermal shock testing, thermal functional testing, calibration, system level testing and on orbit commissioning. A new calibration test process was developed, along with supporting structure and software to ease the testing process, producing accurate calibration parameters and expected performance results for the sensors. A thermal qualification campaign was completed, demonstrating that sun sensors are capable of functioning with negligible performance degradation after exposure to extreme temperatures. A process for installing the sun sensor pin hole was developed using photolithography. Finally, power subsystem analysis for the NEMO-AM mission is presented.

This thesis describes technologies developed for nanosatellites at the Space Flight Laboratory. A critical ground station component, the Terminal Node Controller, was upgraded in order to support Generic Nanosatellite Bus and future missions. Sun sensor requirements and operation were reviewed, followed by details of the author's work in executing the acceptance testing on these parts, including thermal shock testing, thermal functional testing, calibration, system level testing and on orbit commissioning. A new calibration test process was developed, along with supporting structure and software to ease the testing process, producing accurate calibration parameters and expected performance results for the sensors. A thermal qualification campaign was completed, demonstrating that sun sensors are capable of functioning with negligible performance degradation after exposure to extreme temperatures. A process for installing the sun sensor pin hole was developed using photolithography. Finally, power subsystem analysis for the NEMO-AM mission is presented.

In this thesis, a simulation harness is constructed in Simulink for the purpose of validating the Attitude Determination and Control System (ADCS) on the APTAS student CubeSat in support of the upcoming flight readiness review. The simulation results are used to verify the compliance of a subset of the requirements for the ADCS, detailed in table 1. Calibration of the onboard sun sensor array, which is used to find the sun vector for the attitude determination, is performed using a break-out board of sun sensors tested in a sun simulator. The data gathered from this test is used to model the sun sensor system in the simulation and in flight software. The results show that the sun sensor system is able to find the sun vector with an average error angle of 5.4 degrees, though the error angle may spike up to 18 degrees in operation. It is found that the complete ADCS is able to guide the spacecraft toward the desired nadir-facing attitude, though not with the accuracy specified in the requirements. The spacecraft is able to detumble much better than required. All deficiencies found in the ADCS software have been corrected. These changes arelisted in appendix B. It is concluded that, despite its flaws, the ADCS software is flight ready. / Project APTAS