Descent & Landing Trajectory and Guidance Algorithms with Divert Capabilities for Moon Landing

TL;DR

This work develops a fuel-optimal descent and landing trajectory for ESA's Argonaut lunar lander within a four-phase D&L scheme, incorporating divert capability and terrain/hazard sensing constraints. It combines an off-line multi-phase Optimal Control Problem to obtain a nominal trajectory with $m_p\approx3105\,\mathrm{kg}$ and $\Delta V\approx1898\,\mathrm{m/s}$, and a sub-optimal, easily implementable onboard guidance path using PEG-inspired braking, a pitch-up maneuver, and a polynomial descent with diverts. A Differential Evolution offline optimization tunes the sub-optimal trajectory parameters, achieving propellant penalties below 1% compared to the optimal, while maintaining feasibility under divert scenarios. The results indicate the proposed lightweight onboard guidance approach can meet Argonaut’s mission requirements with reduced verification, validation, and storage burden, making it a viable candidate for Phase B2/C/D development.

Abstract

This paper presents the preliminary design of the descent and landing trajectory of the ESA Argonaut lunar lander. The mission scenario and driving system constraints are presented and accounted for in the design of a fuel-optimal trajectory that includes divert capabilities, as required to achieve a safe landing. A sub-optimal descent and landing trajectory is then presented and computed from the optimal one, and the related on-board guidance algorithms are derived. The proposed end-to-end guidance solution represents an easily implementable alternative to on-board optimization, minimizing the verification & validation effort, computational footprint, and programmatic risk in the development of the related GN&C capabilities. A dedicated off-line optimization process is also outlined, and exploited to optimize the propellant consumption of the sub-optimal trajectory and to ensure the fulfillment of system constraints despite the use of simple algorithms on-board. The sub-optimal trajectory is compared to the optimal baseline, and conclusions are drawn on the applicability of the proposed approach to the Argonaut mission.

Figures (9)

• Figure 1: Hypothetical Argonaut lander configuration 3dmodel.
• Figure 2: Descent and landing scenario.
• Figure 3: Spherical coordinates definition.
• Figure 4: Optimal trajectory -- variable thrust during braking burn. Dotted vertical lines indicate the location of LGA, and VGA waypoints.
• Figure 5: Reachable $\theta\left(t_0\right)$ as function of the thrust level $T$ for $\theta_f = 90$ deg. Notice that 1 degree corresponds to $\approx$30 km of along-track displacement at 30 km of altitude.