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.

Re-entry flights of reusable first or upper stages typically foresee phases in the hypersonic flight regime, characterized by severe aero-thermal loads which could become critical for the most exposed components, like the vehicle forebody or the fin leading edges. These require consequently dedicated thermal protection systems (TPS), whose design generally requires a multi-disciplinary approach. In this framework, a viable solution is the use of high-temperature resistant ceramic matrix composite (CMC) structures, but the implementation of such technology, especially for the manufacturing of complex components and its application in real flight conditions, still presents significant challenges. In this work, the design activities for the CMC-based TPS of the payload forebody of a hypersonic sounding rocket are presented, developed within the framework of the STORT project, whose mission includes in flight demonstration of multiple critical technologies required for sustained flight at Mach numbers above 8, corresponding to a significantly high integral thermal load.

The long duration hypersonic flight experiment ATHEAt was carried out with a two-stage sounding rocket configuration from the Andøya Space launch site on 6th October 2025 in Norway. The first stage motor Red Kite combined with the Black Brant Mark IV motor as the second stage made it possible to fly a payload with a mass of 244 kg at Mach numbers between 5.0 and 9.46 at an altitude range from 18 to 31 km for 119 s. The forebody of the scientific payload features a thermal protection system (TPS) made of ceramic matrix composite (CMC) segments. The first part has an ogive shape followed by an octagonal shape and four flaps. The payload includes a transpiration cooling and an impingement cooling experiment. The initial 5 deg deflection of all four flaps was increased to 20 deg during the hypersonic flight phase. In addition to 370 intrusive sensors, additional nonintrusive techniques were used for the measurement of displacement and deformation of hot structures. Two infrared (IR) cameras and four pyrometers measured the surface temperature distribution of the flap rear surfaces. To minimize the apogee dispersion and enable precise science conditions for the flight experiment, a flight-qualified predictor was used.