Lars Witte

Small penetrator probes have been proposed regularly as low(er) cost landing elements, particularly for network science. This study reviews such a concept with regard to its soil penetration capability with numerical and experimental investigations. A reference micro Mars lander (MML) is considered to deliver a 10 kg payload to the Martian surface. It decelerates the penetrator probe with a mechanical decelerator to 40–60 m/s before impact. The remaining kinetic energy is distributed to the soil and an internal load limiter on impact, which restricts the maximum g-load acting on the payload. A semi-empirical force–displacement law and a discrete element method simulation were used to describe penetration behavior. Major impact force constituents are the velocity-dependent, drag-like displacement of the soil particles and the compaction of soil along the penetration path. Based on the simulation results, a test penetrator was designed and tested by impacting it on different soil conditions with a maximum impact velocity of 10 m/s. Cohesive Mars soil simulant, non-cohesive quartz sand, and stones of various sizes were used to assess different impact conditions. The comparison of test and simulation data identifies their respective capabilities and limitations. Recommendations for use and findings for further research are deduced therefrom.

Small penetrator probes have been proposed regularly as low(er) cost landing elements particularly for network science. This study reviews such a concept with regard to its soil penetration capability with numerical and experimental investigations. A reference Micro Mars Lander (MML) is considered to deliver a 10 kg payload to the Martian surface. It decelerates the penetrator probe with a mechanical decelerator to 40–60 m/s before impact. The remaining kinetic energy is distributed to the soil and an internal load limiter on impact, which restricts the maximum g-load acting on the payload. A semi-empirical force-displacement law and a discrete element method simulation were used to describe penetration behavior. Major impact force constituents are the velocity-dependent, drag-like displacement of the soil particles and the compaction of soil along the penetration path. Based on the simulation results, a test penetrator was designed and tested by impacting it on different soil conditions with a maximum impact velocity of 10 m/s. Cohesive Mars soil simulant, non-cohesive quartz sand, and stones of various sizes were used to assess different impact conditions. The comparison of test and simulation data identifies their respective capabilities and limitations. Recommendations for use and findings for further research are deduced therefrom.

Several programmes are dedicated worldwide to the development of reusable vertical take-off vertical landing (VTVL) launch vehicles. Thereby, the touchdown event with its associated impact forces and shocks poses load cases to the vehicle being new in the launcher domain. Furthermore, the application of related knowledge from planetary, exploration-type landers has to be taken cautiously due to the differences between those types of vehicle, particularly with regard to their tank configuration and structural sizing. Depending on flight test conditions and operational concept, non-neglectable amounts of liquid propellant can still be inside the tanks at landing and interact with the vehicle. These interactions cause dynamic and structural effects that affect the landing stability and structural integrity of the vehicle. This study aims to experimentally investigate the fluid–vehicle interactions of a fully functional touchdown demonstrator during touchdown. For this, a vehicle landing engineering model is equipped with a circular cylindrical tank and a series of touchdown tests with varying horizontal landing velocities and fill levels is conducted. Thereby, the main objectives are to prove repeatable landing behaviour under sloshing impact for constant landing conditions, to characterise the fluid impact on landing stability and to investigate the fluid–structure interactions on the tank. Therefore, test data of different test cases are compared and analysed with regard to dynamic behaviour and structural responses.

The objectives for celestial body exploration especially regarding lunar exploration became more and more demanding over recent years. New landers are needed for missions in the Artemis programme by the National Aeronautics and Space Administration (NASA) or the Terra Novae exploration programme by the European Space Agency (ESA) that offer a great deal of reliability and versatility. A system that can level, align, and lower the whole landing platform could offer a lot of benefits in comparison to older lander concepts without such a system. Challenges like unevenness of the ground, slopes in the area of the landing site, as well as different compression of the lander legs can be counteracted. Additionally, access to the lander’s payload bay is simplified by lowering the landing platform as a whole closer to the ground. This study takes a closer look at an approach for the conceptual design of such a mechanism in the early design stages of a lander using a minimal number of known parameters and low computational expenditure. The approach of calculating the motion of the lander and the forces due to the motion on each of the legs primary strut interfaces is first explained and then demonstrated for the European Large Logistics Lander (EL3) on the basis of a mechanism, which moves the primary interface of each landing leg in the vertical axis of the lander.

The Mars Exploration Program Analysis Group has identified measurements of the state and the variability of the Martian atmosphere as high priority investigations for the upcoming years. Balloon-borne instruments could bridge the gap in both temporal and spatial resolution in mesoscale distances between local, stationary landers and global orbiter observations. The idea to use a balloon system for such a purpose is not new in essence and has been proposed already in past decades. While those concepts considered an aerial deployment during entry and descent, the concept outlined in this study revisits a launch off the payload deck of a lander from the Martian surface. This deployment option profits today mainly from the technological advance in micro-electronics and sensor miniaturization, which enables the design of a balloon-probe significantly smaller than earlier proposed systems. This paper presents the feasibility assessment for this instrument and gives further details on the scientific and operational concept, a strawman sensor suite, its system components and the associated size and budget estimates. It is complemented by the analysis scheme proposed to assess, manage and mitigate the deployment risk involved in automatically launching such a balloon-system off a planetary surface.

Prior to landing of reusable space transportation systems, the vehicle’s landing legs needs to be fully deployed to enable a safe landing and further re-use of the space vehicle. During that phase the deployment system has to overcome harsh and challenging environmental conditions. In this study, a numerical simulator is developed in order to investigate these influences on the landing leg deployment dynamics. By means of an extensive aerodynamic database and a broad approach flight domain, the influence of aerodynamics, exhaust plume, and vehicle’s attitude on the deployment dynamics is analyzed. This study shows on the example of the first stage demonstrator CALLISTO (Cooperative Action Leading to Launcher Innovation in Stage Toss back Operations), that thrust level, vehicle attitude, and the deployment system parameters affect the deployment performance.