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Introduction: Apollo planned for the future, retaining a suite of specially curated pristine samples. One of these samples is an Apollo 17 double drive-tube core (73001/73002) [1], which samples down to 70 cm below the lunar surface within the ‘light mantle’ unit in the Taurus Littrow Valley (TLV) [2]. Models estimate a low temperature of ~250 K [3] at this depth, conducive for cold-trapping of volatiles. The lower core segment (73001) has been kept within a Core Sample Vacuum Container (CSVC) since 1972 when it was sealed under vacuum at the surface of the Moon [4]. On Earth, CSVC 73001 was sealed in a secondary outer vacuum container (OVC) at a pressure of ~6×10-2 mbar [5]. One of the science goals for Apollo 17 was to target samples that may have trapped gases released from the lunar interior via the Lee-Lincoln fault [1]. The vacuum sealed CSVC 73001 presents a prime opportunity to investigate for such gases. Here we present the preparation and execution of a unique gas extraction event of CSVC 73001, as part of the Apollo Next Generation Sample Analyses (ANGSA) Program. This activity was led by an ANGSA subteam including the European Space Agency (ESA), Washington University St Louis (WUStL), JSC Apollo Sample Curation Facility and University of New Mexico (UNM). Gas Extraction Set-Up & Challenges: A gas extraction set-up was created composed of two main hardware components: a piercing tool, designed and built by ESA, which interfaces with an ultra-high vacuum (UHV, order of pressure, ×10-9 mbar) gas extraction manifold developed and built by WUStL [6]. Technical and scientific challenges for the hardware design include: preserving the pristinity of the regolith and sample gases, avoiding contamination, preventing isotopic fractionation, capability to operate under UHV; accounting for limitations of dexterity when working within a dry N2 glovebox, and delivering on precision piercing of the CSVC stainless steel base without piercing the Teflon cap of the sample holder within. Design and Manufacture: Following a hardware literature review [e.g., 4,7], an ‘agile’ iterative design approach was undertaken. This included: deriving a set of design requirements based on science, curation and technical needs; experimentally deriving the required piercing force and piercing tip dimension; iterative breadboarding; and regular consultation with ANGSA team members. The piercing tool was machined and manufactured at ESA and the extraction manifold built and calibrated at WUStL. Mechanical testing of the piercing tool under ambient conditions aided operational refinement and piercing tip calibration. Subsequent extensive testing under UHV demonstrated efficient, repeatable use of the piercing tool in meeting the objectives and identified challenges and requirements. The tested hardware underwent stringent cleaning and baking (heated in a vacuum oven at 180ºC for 72 hours) prior to being installed at JSC. Outer Vacuum Container Gas Extraction: The OVC containing the CSVC and constituent sample was interfaced directly with the gas extraction manifold and a ‘blank’ sample of the background collected under UHV (order of ×10-9 mbar). A 100 cc ‘test’ aliquot and two gas samples were sequentially collected of the OVC gas, each for a duration of 15 minutes. The OVC gas could be important, if the CSVC may have leaked over the past 50 years. Initial (uncalibrated for system volume) pressures of the OVC sample (order of ×10-2 mbar) are consistent with that of the OVC when it was originally sealed. Piercing the CSVC: The CSVC was extracted from the OVC and rendered XCT images showed the most challenging scenario of the sample holder with the Teflon cap in direct contact with the CSVC base. The CSVC was transferred to the piercing tool and interfaced with the gas extraction manifold (Fig.1). Extensive He-leak testing indicated no atmospheric leaks. Monitoring of pressures and RGA spectra (using a quadrupole mass spectrometer) indicated that the CSVC may be leaking. The manifold was isolated with the system pressure at 8.7×10-9 mbar and the piercing commenced. A satisfying ‘pop’ marked successful piercing of the CSVC base. Test aliquots (100 cc); two ‘short’ 15 minute duration samples (uncalibrated pressure, ~6 mbar) and ‘long’ extraction (~1-2 week duration) samples were collected. Pressures and RGA spectra were monitored throughout from which a set of hypotheses on how well the OVC and the CSVC have performed have been made and will be tested as part of laboratory analyses of the gas. Analyses include determining general gas composition (Z. Sharp at UNM) and noble gas isotope ratios (R. Parai at WUStL). In particular, 20Ne/22Ne and 128Xe/130Xe can assess for presence of lunar gas [8]. Findings and Lessons Learned: The full performance of the piercing tool and CSVC is still to be fully assessed. First XCT and optical images of the CSVC base indicate a high-level of tool performance in terms of not having pierced the Teflon cap and producing an adequate size of hole. A baseline set of requirements and sequence of events have been demonstrated for a gas extraction procedure, which is informing a new generation of (volatile-rich) sample return missions (e.g., Artemis; Mars Sample Return). Synthesis of the CSVC performance will also inform development of easy to use containers with longer lasting, contaminant-free seals. Other lessons learned include recommendations to: produce physical duplicates and digital twins of flight hardware; further research into hardware degradation during flight and with time; ensure early definition of mission goals and consultation with science teams and curators (critical for providing clear driving requirements for hardware development); include flexibility in sample container designs and interfaces that account for the entire return sample life-time needs and not just at point of collection. Figure 1: Gas extraction setup composed of an UHV manifold and piercing tool. References: [1] Meyer C. (2011) Lunar Sample Compendium. [2] Schmitt et al. (2017) Icarus, 298, 2-33. [3] Keimh S.J. and Langseth M.G. (1973) Proc. 4th Lunar Sci. Conf. 2503-2513. [4] Allton J.A. (1989) JSC23454, NASA. [5] Butler P. (1973) Lunar Sample Info. Catalog, NASA JSC. [6] Parai. R et al. (2021) LPSC LII, Abs #2665. [7] NASA JSC (1971) CSVC Technical Drawing, M-11306. [8] Curran, N.M. et al. (2020) PSS, 182, 104823.

The capability to manufacture items in space is an exploration enabling advancement, and will be crucial for sustainable human exploration as we progress beyond Earth orbit. The extrusion based Fused Filament Fabrication (FFF) method using thermoplastics represents a robust and simple methodology applicable to printing parts for both current and future human spaceflight exploration missions. Understanding the performance and behaviour of the FFF process under varying gravity loads is therefore an important knowledge gap that needs to be addressed in order to fully appreciate the characteristics of space manufactured elements. At present, it is not well understood how gravity can influence the characteristics of such elements fabricated in variable gravity environments. In this study, we detail an experiment conducted on a parabolic flight campaign (PFC) wherein we produced a number of FFF polylactic acid (PLA) polymer test articles and compared them to terrestrially fabricated articles. We report on the methodology and the operational parameters used, as well as presenting an analysis of the samples via optical microscopy and tomography. Compressive, tensile and other technical properties are reported herein. A number of explanations are presented to explain the variance in specimens relative to terrestrial reference samples.

Sandwich structures based on C/C–SiC composites (carbon fibre-reinforced carbon with silicon carbide matrix), manufactured using the liquid silicon infiltration (LSI) process and an in situ joining method, offer high specific stiffness and strength, low thermal expansion, high environmental stability, and temperature resistance. Potential application areas are thermal protection systems (TPS) for spacecraft, optical benches in satellites, and hot structures in aerospace. In this work, C/C–SiC sandwich parts of two different geometries, small sandwich samples and large sandwich structures, were manufactured and tested. Carbon fibre-reinforced polymer (CFRP) plates for the skin panels as well as for the cores were made via warm pressing of prepregs based on a 2D carbon (C) fibre fabric, preimpregnated with phenolic resin. After pyrolysis, carbon fibre-reinforced carbon (C/C) core structures were built up and joined to C/C skin panels. Finally, the resulting C/C sandwich preforms were infiltrated with molten silicon (Si), building up a silicon carbide (SiC) matrix. The resulting C/C–SiC sandwich parts were tested in four-point and three-point bending. The applied forces and the correspondent displacements and strain of the skin panels were determined. The bending and shear stiffness as well as bending moment were evaluated through analytical and finite-element (FE) simulation approaches. Furthermore, failure modes of the sandwich samples were analysed. Sandwich stiffness and ultimate bending moment obtained in the bending tests were close to the expected theoretical values, calculated on the basis of the material properties and the sandwich geometry.

Tensile and fatigue tests were performed on an AA2198 aluminum alloy in the T851 condition in ambient air and liquid hydrogen (LH2). All fatigue tests were performed under load control at a frequency of 20 Hz and a stress ratio of R=0.1. The Gecks-Och-Function [1] was fitted on the measured cyclic lifetimes. The tensile strength in LH2 was measured to be 46 % higher compared to the value determined at ambient conditions and the fatigue limit was increased by approximately 60 %. Both S-N curves show a distinct S-shape but also significant differences. Under LH2 environment the transition from LCF- to HCF-region as well as the transition to the fatigue limit is shifted to higher cyclic lifetimes compared to ambient test results. The investigation of the crack surfaces showed distinct differences between ambient and LH2 conditions. These observed differences are important factors in the fatigue behavior change.

Carbon fiber reinforced carbon-silicon carbide (C/C-SiC) sandwich structures have been developed using the Liquid Silicon Infiltration process and the in situ joining method. They offer high mass-specific stiffness, low thermal expansion, and high environmental stability. Potential application areas are highly precise satellite structures, like optical benches. In this study, sandwich samples were manufactured using prepregs based on 2D carbon fibre fabrics and a phenolic resin precursor. Carbon fibre reinforced polymer preforms for folded and grid-cores, as well as for the skin panels were manufactured using autoclave technique. In the second step, the sandwich components were pyrolyzed, leading to C/C preforms. For the build-up of the sandwich samples, two skin panels were joined to a core structure and subsequently, the resulting C/C sandwich preform was siliconized. C/C-SiC sandwich samples were tested under shear load. Shear strength, modulus, and fracture strain were determined and compared to the results obtained by analytical calculation. The shear properties were dependent on the fiber orientation in the core structure as well as on the core type and orientation. The sandwich shear stiffness obtained in the tests was close to the expected theoretical values, calculated on the basis of the material properties and the core geometry.

Cold welding refers to an effect related to space (vacuum). The heavy vibrations during a launch subject interfaces (hold down points) to oscillating motions which may lead to formation of a kind of “friction weld”. If so, these mechanisms may get stuck, and deployment will be hindered. This may endanger the functionality of the mission (instruments) or even the whole spacecraft (if solar panels do not open). Several studies have been done to characterize material combinations (including coatings) for their ability to cold welding in space. Meanwhile, also during launch grease free contacts are demanded. If grease hat to be omitted, the risk of cold welding under fretting was found to increase (when testing in high vacuum). To rate this risk under launch conditions, the test method was recently extended for testing under launch conditions. The new tests procedure consists of fretting applied in the sequence in air, low vacuum and high vacuum. The paper shall present first results gained with this new method of testing in launch conditions and compare them to previous studies done in vacuum. Following the need of space industry on mechanisms for launch and in-orbit life, a first set of combinations of materials and coatings were selected for this new test sequence where fretting is now applied in a sequence of air, low vacuum and high vacuum. Under this sequence, the measured levels of adhesion and it’s evolvement was found to differ strongly from tests done formerly. The paper outlines these first results and compares them to existing data.