Small satellites with increased capabilities in terms of power and propulsion are being demanded for future missions. This paper proposes a possible solution which is the design of a novel integrated solar thermal system that co-generates propulsion and power on-board mini satellites. The system consists of a solar thermal propulsion system (STP) coupled with a micro-Organic Rankine Cycle (ORC) system to harness the waste heat from the STP receiver to provide electrical power and mitigate the need for solar panels. STP provides an alternative to conventional propulsion systems for missions requiring velocity changes of between 800 m/s and 2500 m/s. Additional advantages include higher specific impulses than chemical propulsion systems, throttability, re-start capabilities, and faster transfer times than electrical propulsion systems. The faster transfer times are especially useful for missions that travel across high radiation regions such as the Van Allen Belt. This unique configuration shares resources such as the concentrator and receiver to potentially extend the power and propulsion capabilities while adhering to the strict mass and volume constraints of small satellites. However, there is currently no literature available on the design process of the proposed bi-modal system. This paper therefore presents an integrated solar thermal design strategy for a Geostationary Transfer Orbit to Lunar orbit insertion mission. The design methodology is described in detail to assist with future evaluations of integrated solar thermal systems for other applications and missions. The system is designed to provide a velocity increment of 1.6 km/s. Five mini-satellite sizes were investigated with a gross wet mass of 100 kg, 200 kg, 300 kg, 400 kg, and 500 kg respectively. Each satellite requires to produce an electrical power of 1 W/kg. The STP system uses water as the propellant due to its safety and performance attributes. Toluene has been selected as the working fluid for the ORC due to its high thermal efficiency. By incorporating the use of a high-temperature receiver, propellant temperatures around 2500 K can be achieved that can produce high specific impulse values of more than 300 s. The design has been optimized for various design parameters, such as propellant temperature, nozzle area ratio, burn time, concentrator design, and ORC cycle pressures. The optimization provides an initial framework in the selection of an optimal integrated solar thermal design for the proposed Lunar mission. An analysis of variance has also been conducted to identify which system parameters, such as optical efficiency and turbine efficiency, have the most influential effect on the system. The heaviest components of the system are the propellant (40 to 50%), concentrator (8%), and insulation (8%) with respect to the gross mass of the satellite.