Preliminary Design of a Stand-Alone Mars CubeSate Mission Integrating DLR In-House Technologies

Kurzfassung

Recent progress in CubeSat technology has allowed miniaturized satellites to venture away from Earth’s orbit. In 2018, NASA’s MarCO mission performed a Mars fly-by to provide relay communications between the InSight lander and Earth. In 2022, LICIACube carried out observational analysis of the Didymos asteroid binary system after DART’s impact on Dimorphos. More recently, ten 6U CubeSats built by universities and research centres were launched as secondary payloads of the Artemis 1 mission. These missions will independently explore cislunar and interplanetary space with the lowest cost up to date. Thus, the next significant advancement should encompass a dedicated CubeSat mission to explore a planet in Earth’s proximity. To contribute to this goal, the Institute of Space Systems of the German Aerospace Center (DLR) has developed radiation-hardened small satellite technologies, including communications, power and onboard computer subsystems. This study presents a stand-alone Mars exploration mission using the CubeSat standard that will demonstrate DLR’s in-house technologies. The ConOps is planned for a 4-year mission. After deployment, the spacecraft will carry out a stand-alone Earth-Mars transfer. Upon arrival at Mars, the spacecraft will execute a low-thrust orbital insertion into a highly elliptical orbit. Once in orbit around the Red Planet, the spacecraft will perform aerobraking to position itself around a Primary Science Orbit (PSO) in the altitude range of 250-200 km, where the science operations will take place. This PSO will be a frozen and Sun-synchronous orbit in order to take advantage of Mars’ gravitational parameters that enable a near-circular orbit [32]. During this phase, the spacecraft will tackle two of the most relevant scientific objectives around Mars [14]: gathering data on its lower atmosphere and gravity field. Finally, the spacecraft will be placed into a quarantine orbit for disposal. A system concept is created to accomplish this mission by integrating the in-house technologies, investigating the necessary Commercial-Off-The-Shelf (COTS) components, and performing a feasibility assessment. The resulting 12U CubeSat has a 20.5 kg wet mass, 6.3 km/s manoeuvring capability, and can generate up to 90 W of power at Mars. Its main payloads are a 2U infrared spectrometer, a 1U gravimeter and a 12 Mpx CMOS camera. The propulsion system is a Busek BIT-3 gridded-ion engine as used in the Lunar IceCube and LunaH-Map missions [45]. This system features an expanded tank to accommodate the needed 5 kg of propellant. The XACT-50 developed by Blue Canyon Technologies is used as the integrated attitude control and determination system. The DLR Integrated Core Avionics stack will be used, which combines multiple avionics systems (i.e., onboard computer, communications, and power) into a single unit. The DLR ScOSA onboard computer will be used, which integrates two computing nodes to ensure robustness. The reliable computing node comprises the OBC used on the MASCOT lander developed by DLR. The high-performance computing node consists of a COTS-based processing module for application acceleration. The spacecraft uses the radiofrequency X-band in order to communicate with its ground segment. The in-house developed Generic Software Defined Radio (GSDR) is used as a transceiver. The system also features a main reflectarray antenna as used by MarCO [29] and a secondary patch antenna. Electrical power is generated using two linear solar arrays with 32% effciency triple junction solar cells developed by Azur Space. The power control and distribution unit, as well as the batteries, will be adapted from those used by the DLR PLUTO mission. Thermal coating and heaters are used for thermal management; a preliminary thermal analysis is presented that demonstrates their feasibility. All of these systems are housed in an EnduroSat 12U XL main structure. A ground segment architecture is presented using a combination of DLR’s antennas, ESA’s ESTRACK infrastructure and NASA’s Deep Space Network. These are also used for navigation purposes using ranging, Doppler and delta-DOR measurements. Different launch opportunities are studied, favouring a piggyback on another Mars-bound launch. A preliminary mission analysis is performed using the TU Delft Astrodynamics Toolkit (Tudat) to assess the feasibility of the mission. Assuming the launch method mentioned above, a set of porkchop plots is generated to constrain the possible launch windows. The hodographic shaping method [26] is then used to effciently calculate the available low-thrust Earth-Mars trajectories. An optimisation process is implemented using a grid search over the potential departure dates and times of flight for these trajectories. The behaviour of the optimal-Delta-V hodographic trajectories across this grid is evaluated using a differential evolution algorithm. An optima region has been found for the studied departure dates with TOFs in the range of 1000 to 1400 days. A 25% increase over the first estimate of Delta V for the mission enables feasible trajectories for 80% of the departure dates studied, with room for improvement. These trajectories exhibit maximum thrust values in the same order of magnitude as the system requirements, which is encouraging for a first estimation. An orbit insertion strategy is proposed involving a ballistic capture. Finally, a study of a Mars science orbit that enables the scientific objectives of the mission is presented.