Session
Technical Session V: Standards and Modularity
Abstract
Spacecraft standardization has been a topic of great debate within the space community. This paper intends to be a provocative thought piece asking one fundamental question: “is there a ‘right size’ for small satellites?” In order to answer this question, we propose three top-down design factors for the space systems engineering process: spacecraft utility, mission utility, and optimum cost. Spacecraft utility quantitatively measures the capability of a spacecraft, derived from its volume and power properties. Mission utility then measures the aggregate value of a constellation. Optimum cost, which is a function of spacecraft mass and quantity, can be determined by assessing the break-even point. Data from the small satellite community, including USAF Academy FalconSAT and Surrey Satellite Technology Ltd. (SSTL) missions, is presented in support of this discussion, constrained to systems with a mass less than 200 kg. These design factors inform the mission developer in determining the appropriate system architecture. Using these design factors, a notional standardized spacecraft configuration is presented, with a mass of 30 kg and 50 cm cubed volume that optimizes spacecraft utility, mission utility, and cost.
Presentation
Right-sizing Small Satellites
Spacecraft standardization has been a topic of great debate within the space community. This paper intends to be a provocative thought piece asking one fundamental question: “is there a ‘right size’ for small satellites?” In order to answer this question, we propose three top-down design factors for the space systems engineering process: spacecraft utility, mission utility, and optimum cost. Spacecraft utility quantitatively measures the capability of a spacecraft, derived from its volume and power properties. Mission utility then measures the aggregate value of a constellation. Optimum cost, which is a function of spacecraft mass and quantity, can be determined by assessing the break-even point. Data from the small satellite community, including USAF Academy FalconSAT and Surrey Satellite Technology Ltd. (SSTL) missions, is presented in support of this discussion, constrained to systems with a mass less than 200 kg. These design factors inform the mission developer in determining the appropriate system architecture. Using these design factors, a notional standardized spacecraft configuration is presented, with a mass of 30 kg and 50 cm cubed volume that optimizes spacecraft utility, mission utility, and cost.
