Martin Frieß

Increasing the efficiency of jet engines is essential to meet the demanded climate targets. Ceramic matrix composites (CMCs) are strong candidates for aircraft applications because they withstand high temperatures, while their density is two-thirds lower than that of conventional nickel-based alloys. This leads to cooling air savings and a lower overall engine weight, resulting in a potential reduction of emissions. To investigate the potential benefits and manufacturing techniques required for the introduction of CMC to the high-pressure turbine (HPT) of a modern jet engine, the geometry of a nozzle guide vane of an existing turbine was redesigned considering ceramic specific constraints. Then, the liquid silicon infiltration (LSI) process was used to manufacture silicon carbide fiber-reinforced silicon carbide (SiC/SiC) nozzle guide vanes (NGV). Hi-Nicalon S woven fabric was used together with a chemical vapor infiltration (CVI)-based fiber coating. The outer surface of the vane was ground to meet the requirements for surface roughness, and geometric and positional tolerances. Cylindrical, laser-drilled cooling holes were introduced for trailing edge cooling. In the final step, an environmental barrier coating (EBC) system consisting of yttrium disilicate (Y-DS) and yttrium monosilicate (Y-MS) layers was applied using physical vapor deposition (PVD) processing. Wind tunnel testing under technology readiness level (TRL) 4 will be performed and vane performance will be evaluated.

In the ongoing development of hypersonic technologies, material advancements play a key role in meeting the ever-increasing thermomechanical demands of these applications. Ultra-High Temperature Ceramic Matrix Composites (UHTCMCs) offer a promising solution for components operating under such extreme conditions. Their outstanding thermomechanical properties, including high temperature and thermal shock resistance, excellent thermal conductivity and mechanical strength, position them as ideal candidates for applications in fields like leading edges or inlet ramps for ramjets and scramjets. Due to their remarkable material composition, UHTCMCs are capable of operating in temperature regimes that surpass 1700 °C during their operation times under oxidizing atmospheres. At the German Aerospace Center (DLR), a UHTCMC material based on carbon fibres and a zirconium diboride matrix is being developed utilizing Reactive Melt Infiltration (RMI). With RMI, the orientation of the reinforcement fibres can be tailored, to enable the material to fulfill the demanding load requirements. The reactive melt infiltration process comprises three stages: preform fabrication, pyrolysis, and the actual melt infiltration. The foundation for important material properties of the final ceramic, including the matrix composition, is established in the preform production, which is a crucial step in the process. A boron- and zirconium diboride-based slurry is infiltrated into pitch-based carbon fibre fabric. Subsequently, the preforms are consolidated, pyrolysed, and infiltrated with molten Zr₂Cu to obtain the UHTC matrix by in situ reaction with the preform elements. Scanning Electron Microscopy (SEM) and Energy-dispersive X-Ray Spectroscopy (EDX) enable the examination of the microstructural features, including the arrangement and distribution of zirconium diboride within the matrix. Mechanical evaluation of the UHTCMCs is conducted via 3-point bending tests at both room temperature and at elevated temperature at 900 °C. It has been demonstrated that Ultra-High Temperature Ceramic Matrix Composites can be produced by means of reactive melt infiltration, and that they retain their strength even at elevated temperatures.

For the development of advanced materials and structures with improved corrosion and abrasion resistance, a better understanding of the complex multiphase flow/structure interaction is necessary. Besides appropriate material development of ceramic matrix composites (CMC) materials concerning the harsh application conditions, a precise material characterization with respect to compatibility to relevant fuels on aging aspects and, especially, to performance on application-related combustion tests for ramjet and scramjet propulsion are of special interest. In order to test and evaluate different silicon carbide (SiC) contents on the surface of different C/C-SiC materials, two different materials of quality XS (standard) and an enhanced C/C-SiC material with SiC enriched surface and graded structural lay-up were developed, manufactured, characterized, and tested at DLR's test facility M11 at scramjet conditions. This facility is able to produce the boundary conditions typical of a scramjet/ramjet propelled flight at Mach 5.5 to Mach 6.0 flight at an altitude of 30 km in the connected pipe ground test configuration. It is capable of stagnation temperatures up to 1500 K, 25 bar stagnation pressure, and 5 kg/s hot gas mass flow rate. The material samples were exposed to the vitiated air in a scramjet model combustion chamber with lateral optical access and a wedge-shaped shock generator similar to flame holder geometries This publication gives both a detailed overview of the material and component development and an overview of the hot gas testing facility and its capabilities for testing materials and components. In addition, some typical results of tested C/C-SiC materials are presented.

The paper presents manufacture of C/C‐SiC composite materials by wet filament winding of C fibers with a water‐based phenolic resin with subsequent curing via autoclave as well as pyrolysis and liquid silicon infiltration (LSI). Almost dense C/C‐SiC composite materials with different winding angles ranging from ±15° to ±75° could be obtained with porosities lower than 3% and densities in the range of 2 g/cm ³ . Thermomechanical characterization via tensile testing at room temperature and at 1300°C revealed higher tensile strength at elevated temperature than at room temperature. Thus, C/C‐SiC material obtained by wet filament winding and LSI‐processing has excellent high‐temperature strength for high‐temperature applications. Crack patterns during pyrolysis, microstructure after siliconization, and tensile strength strongly depend on the fiber/matrix interface strength and winding angle. Moreover, calculation tools for composites, such as classical laminate and inverse laminate theory, can be applied for structural evaluation and prediction of mechanical performance of C/C‐SiC structures.

Oxide ceramics could be attractive high-temperature construction materials for critical structural parts in high-temperature molten salt thermal energy storage systems due to their excellent corrosion resistance and good mechanical properties. The 3D-printing technology allows the production of ceramic components with highly complex geometries, and therefore extends their applications. In this work, 3D-printed ZrO₂ and Al₂O₃ ceramics were immersed in molten MgCl₂/KCl/NaCl under argon or exposed in argon without molten chlorides at 700 °C for 600 h. Their material properties and microstructure were investigated through three-point-bend (3PB) testing and material analysis with SEM-EDX and XRD. The results show that the 3D-printed Al₂O₃ maintained its mechanical property after exposure in the strongly corrosive molten chloride salt. The 3D-printed ZrO₂ had an enhanced 3PB strength after molten salt exposure, whereas no change was observed after exposure in argon at 700 °C. The material analysis shows that some of the ZrO₂ on the sample surface changed its crystal structure and shape (T→M phase transformation) after molten salt exposure, which could be the reason for the enhanced 3PB strength. The thermodynamic calculation shows that the T→M transformation could be caused by the reaction of the Y₂O₃-stabilized ZrO₂ with MgCl₂ (mainly Y₂O₃ and ZrO₂ with gaseous MgCl₂). In conclusion, the 3D-printed ZrO₂ and Al₂O₃ ceramics have excellent compatibility with corrosive molten chlorides at high temperatures and thus show a sound application potential as construction materials for molten chlorides.