Original approaches based on ablative materials and novel TPS solutions are required for space applications where resistance in extreme oxidative environments and high temperatures are required. The atmospheric entry of space vehicles from high-energy trajectories requires high-performance thermal protection systems that can withstand extreme heat loads.
A new scenario has appeared due to a worldwide change in space mission planning strategies with entry vehicles going back to capsule designs and ablators are re-gaining attention. Consequently, the development of new, more efficient materials and systems is a must. Such developments, nevertheless, have to be subject to extensive experimental investigations using suitable facilities to obtain the desired maturity level and optimization. In this view, the investigation and development of new materials based on ablative and ceramic thermo-structural concepts is crucial. A new (hybrid) concept based on the combination of both type of TPS materials was proposed. The advantage of the ceramic for this function is the low density compared to ablative material and the excellent thermal performance in this heat load range, as well as the stability of the shape of TPS which is an advantage for the aerodynamic of the re-entry vehicle.
The concept of the project is based on the development of a novel hybrid heat shield, based on the integration of an external ablative part with a ceramic matrix composite (CMC) thermostructural core. This has been carried out by the integration of dissimilar materials. The main advantage of a hybrid TPS heat shield is based on the capability of the ablative layer of the hybrid TPS of bearing higher heat loads than the ceramic layer underneath while the tough ceramic composite underneath provides structural support. There was a big challenge to achieve a sound bonding between the two parts. This was carried-out by employing advanced bonding technologies. The development of new adhesives solutions with improved mechanical and insulating characteristics was investigated. The use of advanced high temperature adhesives and hybrid solutions in combination with mechanical attachments was assessed, as well as other existing hybrid solutions. From this point of view the HYDRA system is offering improved mechanical properties as well as better robustness during the entry. Besides, the new moon or interplanetary planned missions create higher heat loads during earth re-entry than ceramic or metallic TPS can withstand. Since these heat loads are characterized by a peak profile the ablator can dissipate the high heat loads during the peak. For that, a comparatively thin layer of ablative material is thought to be sufficient.
The HYDRA project started at the beginning of 2012 and had duration of three years. The core group of the project was composed of 10 public and private organisations giving an excellent balance between large industries (Airbus DS GmbH and Airbus DS SAS), SMEs (HPK and HPS - Portugal), Public research entities and Universities (NCSRD, INCAS, ICMCB and IRS) and private research centres (TECNALIA, DLR). The partners are coming from five different European countries: France, Greece, Germany, Romania and Spain. The project consisted of seven different technical work packages dealing with: the selection of a reference mission and specifications, definition of the current state-of-the-art and materials trade-off, procurement of ablator and CMC parts, study of the ablator to CMC attachment, simulation & design and characterisation at relevant environment to achieve a TRL 4 at the end of the project.