This study focused on the thermal stability and ablative behavior assessment of five newly developed composite TPS configurations. All ten test samples were 3D printed via FDM using various fire-retardant thermoplastic materials, with and without reinforcement. Eight samples integrated a new thermal management internal air chamber conceptualized architecture. A prompt feasible approach for the flame resistance preliminary assessment of ablative TPS samples was developed, using an in-house oxy-acetylene torch test bench. Experimental OTB ablation tests involved exposing the front surface samples to direct flame at 1450 ± 50 °C at 100 mm distance. For each configuration, two samples were tested: one subjected to 30 s of flame exposure and the other to 60 s. During testing, internal temperatures were measured at two backside sample contact points. Recorded temperatures remained below 46 °C, significantly under the maximum allowable back face temperature of 180 °C set for TPSs. The highest mass losses were measured for PC and PETG FR materials, achieving around 19% (30 s) and, respectively, 36% (60 s), while the reinforced configurations showed overall only a third of this reduction. The study’s major outcomes were the internal air chamber concept validation and identifying two reinforced configurations as strong candidates for the further development of 3D-printed ablative TPSs.

The research investigated the potential of five novel additively manufactured (AM) fiber-reinforced thermoplastic composite (FRTPC) configurations as alternatives for ablative thermal protection system (TPS) applications. The thermal stability and ablative behavior of ten samples developed via fused deposition modeling (FDM) three-dimensional (3D) printing out of fire-retardant thermoplastics were investigated using an in-house oxyacetylene torch bench. All samples featured an innovative internal thermal management architecture with three air chambers. Furthermore, the enhancement of thermal benefits was achieved through several approaches: ceramic coating, mechanical hybridization, or continuous fiber reinforcement. For each configuration, two samples were exposed to flame at 1450 ± 50 °C for 30 s and 60 s, respectively, with the front surface subjected to direct exposure at a distance of 100 mm during the ablation tests. Internal temperatures recorded at two back-side contact points remained below 50 °C, well under the 180 °C maximum allowable back-face temperature for TPS during testing. Continuous reinforced configurations 4 and 5 displayed higher thermal stability the lowest values in terms of thickness, mass loss, and recession rates. Both configurations showed half of the weight losses measured for the other tested configurations, ranging from approximately 5% (30 s) to 10–12% (60 s), confirming the trend observed in the thickness loss measurements. However, continuous glass-reinforced configuration 5 exhibited the lowest weight loss values for both exposure durations, benefiting from its non-combustible nature, low thermal conductivity, and high abrasion resistance intrinsic characteristics. In particular, the Al₂O₃ surface coated configuration 1 showed a mass loss comparable to reinforced configurations, indicating that an enhanced surface coat adhesion could provide a potential benefit. A key outcome of the study was the synergistic effect of the novel air chamber architecture, which reduces thermal conductivity by forming small internal air pockets, combined with the continuous front-wall fiber reinforcement functioning as a thermal and abrasion barrier. This remains a central focus for future research and optimization.

This study investigates seven advanced hybrid composite thermal protection system (TPS) prototypes, featuring an innovative internal air chamber design that reduces heat conduction and enhances overall thermal protection performance. Specimens were manufactured by fused deposition modeling (FDM), an additive manufacturing technique, using a fire-retardant thermoplastic. Selected configurations were reinforced with continuous carbon or glass fibers, coated with ceramic surface layer, or hybridized with carbon fiber reinforced polymer (CFRP) layers or a CFRP laminate disk. To validate performance, a harsh oxy-acetylene torch (OAT) protocol was implemented, deliberately designed to exceed the severity of most reported typical ablative assessments. The exposed surface of each specimen was subjected to direct flame at a 50 mm distance, recording peak temperatures of 1600 ± 50 °C. Two samples of each configuration were tested under 60 and 90 s exposures. Back-face thermal readings at potential payload sites consistently remained below 85 °C, well under the 200 °C maximum standard threshold for TPS applications. Several configurations preserved structural integrity despite the extreme environment. Prototypes 4.1 and 4.2 demonstrate the most favorable performance, maintaining structural integrity and low back-face temperatures despite substantial thickness loss. By contrast, specimen 6.2 exhibited rapid degradation following 60 s of exposure, which served as a rigorous and selective early-stage screening tool for evaluating polymer-based ablative TPS architectures.

Sustained cislunar logistics operations, including recurring support of the Lunar Gateway at EML1 and EML2, impose demanding propulsion requirements, including high ΔV budgets, restart capability, and long-duration propellant storage, which conventional propulsion approaches struggle to meet efficiently at scale. This study presents a novel cislunar mission architecture based on nuclear thermal propulsion (NTP), operating at a specific impulse of 900 s with liquid hydrogen as propellant and a hydrazine Reaction Control System (RCS) for proximity and docking maneuvers. The architecture is evaluated analytically through sequential application of the Tsiolkovsky rocket equation across two mission scenarios: a direct logistics transfer to EML1 (Scenario A) and a two-burn Gateway staging transfer from EML1 to EML2 (Scenario B), using a launch mass of 9000 kg, a 5% ΔV margin, and deterministic ΔV values of 3164 m/s for LEO→EML1, 160 m/s for EML1→EML2, and 37.36 m/s for RCS operations. The proposed architecture achieves a total propellant mass below 3044 kg and a total delivered mass between 5956 kg and 6071 kg across both scenarios. These results establish NTP as a technically credible foundation for scalable and sustainable cislunar transportation, with broad implications for the development of a permanent lunar economy.

Effective thermal insulation of cryogenic liquid hydrogen (LH₂) storage tanks remains a critical engineering challenge, as conventional vacuum-based or monolithic systems are constrained by manufacturing complexity, mechanical vulnerability, and poor geometric adaptability. This study presents the design and numerical verification of a four-layer octagonal composite thermal shield fabricated via additive manufacturing: an AA5083 structural layer (5 mm), a boron nitride-doped ceramic plate (1 mm), up to 290 stacked graphene sheets in a sealed compartment, and an outer Fe₃S₄-TiO₂ nanocomposite layer (~30 µm). Steady-state and transient FEA in ANSYS evaluated three convective boundary conditions (h = 10, 15, and 20 W/m²·K), with the inner wall fixed at 20 K. Temperature distributions remained essentially invariant across all cases (20 K inner, ~20.12 K outer), confirming that thermal performance is governed by the multilayer architecture rather than convective intensity. The shield achieved a mean heat flux of 1684 W/m², R_total ≈ 0.163 m²K/W, and a boil-off rate of 13.9 g/hour. Comparative FEA against NASA US9617069 (q = 193.35 W/m²) and JP2018-119634A (q = 37.975 W/m²) highlights the compactness advantage of the proposed 6 mm shield; the coupled thermo-structural assessment yielded a safety factor of 64,182, confirming elastic-regime operation at 20 K.

The present study investigates the combined influence of manufacturing route and material configuration on the mechanical, elastic, viscoelastic, and impact behavior of Kevlar–carbon hybrid laminates. Three eight-ply laminate configurations (V1, V2, and V3) were manufactured through distinct technological routes: fully prepreg-based hot pressing, Kevlar-prepreg/dry-carbon hand lay-up followed by vacuum curing, and multi-stage hybrid consolidation combining repeated hot pressing with subsequent vacuum curing. The experimental characterization included tensile tests according to ASTM D3039, compression tests according to ASTM D695, determination of Young’s modulus from extensometer measurements and Poisson’s ratio using strain-gauge instrumentation, dynamic mechanical analysis (DMA), and low-velocity impact tests under controlled energy conditions. The novelty of this work consists in the integrated process–configuration–property comparison of these Kevlar–carbon hybrid routes within the same experimental framework, rather than in a generic demonstration that manufacturing affects composite laminates. The V1 laminate exhibited the highest strength-related performance, reaching an average tensile strength of 335.88 MPa and a compressive strength of 165.85 MPa, and it also showed the highest DMA storage modulus at 30 °C, E’ = 53.42 GPa. The V2 laminate presented lower tensile performance but the most pronounced damping response, with the highest tanδ peak value. The Young’s modulus determined from the extensometer measurements was 29.26 ± 1.45 GPa for V1, 26.10 ± 0.22 GPa for V2, and 29.52 ± 1.27 GPa for V3, indicating comparable longitudinal stiffness for the V1 and V3 laminates. The results indicate that the measured behavior is governed by the combined effects of reinforcement form, matrix/resin arrangement, consolidation route, and laminate architecture. Direct quantification of laminate compaction, fiber volume fraction, and void content was outside the scope of the present experimental campaign and is identified as a necessary step for future validation.