The phenomenon by which a spacecraft can both approach a celestial body and start performing revolutions around it, without needing to manoeuvre in between, is known as ballistic capture. For ballistic capture to take place, the spacecraft must be under the gravitational influence of at least 2 celestial bodies. Therefore, it is possible to design heliocentric transfers to Mars culminating in ballistic capture. With an impulsive-thrust strategy, these have already been studied, but were found to be less fuel-efficient and longer-lasting than Hohmann transfers. The objective of the present thesis is to investigate the characteristics of Earth–Mars low-thrust transfers to ballistic capture.

Small spacecraft are very mass- and power-constrained, so orbit transfers are challenging for them, especially to interplanetary destinations. To try and shift this paradigm, the study was carried out assuming the spacecraft to be a 16-unit CubeSat. In addition, to improve the relevance of the results for the design of a real mission, the decision was made to model the spacecraft's environment with many perturbing forces, including third-body perturbations, solar radiation pressure and non-spherical gravity. Furthermore, the performance of the thruster was modelled as being a function of the spacecraft's distance to the Sun.

Ballistic capture orbits at Mars were found in a systematic way, by investigating the past and future behaviour of candidate states with respect to Mars and collecting the states corresponding to capture into a capture set. The most regular orbits of the capture set were found among those making the closest approaches to Mars, but the opposite appears to be true about the orbits with the longer lifetime, which can reach more than a decade. Furthermore, two groups of capture orbits were identified: one coming from the inside of Mars' orbit, the other from the outside, with the orbits of each group being relatively close to each other.

Some capture orbits were selected, each with a different arrival date at Mars, and targeted from Earth, on multiple departure dates. The spacecraft was assumed to leave Earth's orbit with the planet's velocity and the heliocentric transfers were designed with an interior-point method, after direct transcription and collocation. It was found that if the spacecraft is given enough time, the low-thrust strategy requires roughly the same fuel regardless of Earth departure or Mars arrival dates. In addition, terminating a low-thrust transfer to Mars in ballistic capture does not carry additional costs, when compared to simply rendezvousing with the planet. The revolutions that the spacecraft is guaranteed to perform around Mars are then cost-free. With the assumed spacecraft and departure conditions, only around 5 kg of propellant are required to reach Mars and get ballistically captured. Nevertheless, the spacecraft needs to fly for at least 3.5 years, which can be too long for a CubeSat.