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The digitalization of railway systems should increase the efficiency of the train operation to achieve future mobility challenges and climate goals. But this digitalization also comes with several new challenges in providing a secure and reliable train operation. The work resulting in this paper tackles two major challenges. First, there is no single university curriculum combining computer science, railway operation, and certification processes. Second, many railway processes are still manual and without the usage of digital tools and result in static implementations and configurations of the railway infrastructure devices. This case study occurred as part of the Digital Rail Summer School 2021, a university course combining the three mentioned aspects as cooperation of several German universities with partners from the railway industry. It passes through all steps from a digital Control-Command and Signalling (CCS) planning in ProSig 7.3, the transfer, and validation of the planning in the PlanPro data format and toolbox, to the generation of code of an interlocking for the digital CCS planning to contribute to the vision of test automation. This paper contributes the experiences of the case study and a proof-of-concept of the whole lifecycle for the Digital Testfield of Deutsche Bahn in Scheibenberg. This proof-of-concept will be continued in ongoing and following projects to fulfill the vision of test automation and automated launching of new devices., Comment: to be presented at the 1st International Workshop on Testing Distributed Internet of Things Systems

IR pulsed laser radiation in air was applied to Si3N4 and Invar to obtain reliable Si3N4/Si3N4 and Si3N4/Invar adhesive bonded components. The laser pre-treatment produced a homogeneous nanostructured oxide layer on the surfaces, which effectively increased the adhesion at the adhesive/adherends interface and led to cohesive failure in the joining material. The mechanical strength of Si3N4/ Si3N4 and Si3N4/Invar joined components was measured, with and without laser nanostructuring, before and after thermal cycling from room temperature to 50 K, and it resulted that the exposure to extremely low temperatures did not affect the mechanical integrity of the joints. It was also demonstrated that this laser pre-treatment did not alter the mechanical properties of the ceramic substrate.

The effect of a pulsed laser irradiation (Nd:YVO4, 1064 nm) in air on the surface morphology and chemical composition of silicon carbide and on the adhesion with an epoxy adhesive was investigated. Scanning and transmission electron microscopies, atomic force microscopy, and X-ray photoelectron spectroscopy revealed that the laser treatment reduced the contamination level of the surface and induced the formation of a silica-based nanostructured columnar layer on the SiC surface. The mechanism for the formation of five different microstructural regions is described in this paper. In addition, the formation of a 5-10-nm-thick graphite layer between the oxide layer and SiC was observed. The joining test with Hysol® EA9321 showed that the nanostructured columnar silica layer was completely infiltrated by the adhesive, thus leading to a significant increase in the joined specific area and a mechanical interlocking at the adhesive/substrate interface. Nevertheless, the apparent shear strength of the joined SiC samples slightly decreased after the laser processing of the surfaces from about 42 MPa for lapped SiC to about 35 MPa for laser-nanostructured SiC. The formation of the graphite layer was found to be responsible of the poor adhesion properties of the SiC surfaces modified by the laser radiation.

The design of interfaces, coupled with suitable joining materials and joining technologies, is a key parameter in component manufacturing and surface pre-treatments are crucial steps for the production of robust and reliable bonds. In this work, plasma pre-treatments on SiC and Si3N4 are studied and discussed. The pre-treated joined samples have shown a higher mechanical strength than the reference value (lapped surfaces) and the increased adhesion at the ceramic/adhesive interface has led to cohesive failure of the adhesive.

The current work reports a novel approach for the integration of external protective SIC multilayers with ceramic matrix composite (C-f/SiC) with the view of application in aerospace heat protection systems. The integration method is based on diffusion brazing bonding. As a joining agent the MAX-Phase Ti3SiC2, produced by self-propagating high temperature synthesis, has been employed. The pressure applied during the joining process and its effect on the microstructure of the integrated structure is discussed. Microstructural analysis of the resulting joints is conducted using scanning electron microscopy coupled with energy dispersive spectroscopy and X-ray diffraction measurements. Analysis of the joints showed that the bonds are uniform, dense, with few crack vertical to the interface which are not detrimental for the performance of the joints. Ground re-entry tests showed that the joints survive 5 re-entry cycles at 1391 and 1794 degrees C without any detectable damage. (C) 2015 Elsevier Ltd. All rights reserved. European Project "SMARTEES (G.A.) 262749, European Community

In the present study, we elucidate the influence of oxidative heat exposures at 1000 and 1200 °C on an alumina fiber-reinforced polymer-derived ceramic matrix composite containing small residual amounts of carbon. Therefore, we investigated the flexural performance and fracture toughness of on- (0°/90°) and off-axis (45°) reinforced samples. Acoustic emission was used to monitor the internal damage and its progression during loading. At 1000 °C, a moderate reduction of strength and fracture toughness is found while after exposure to 1200 °C a dramatic decrease down to 50 % is observed. For all composites, a reduction of the damaged volume was found after heat treatments indicating a decrease of crack deflection. However, especially at 1000 °C, composites reinforced in 0°/90° direction seemed to be more affected, as no detrimental effect on the mechanical performance was found for the 45° composites. Remarkably, the oxidation-induced silica formation increases the absolute and relative damage thresholds of all composites. A Griffith-like linear relationship between strength and toughness is found. These findings are pivotal for designing and engineering next generation CMCs toward long-term applications.

Reusable thermal protection systems of reentry vehicles are adopted for temperatures ranging between 1000 and 2000 °C, when gas velocity and density are relatively low; they exploit the low thermal conductivity of their constituent materials. This paper presents a new class of light structural thermal protection systems comprised of a load bearing structure made of a macroporous reticulated SiSiC, filled with compacted short alumina/mullite fibers. Their manufacturing process is very simple and does not require special devices or ambient conditions. The produced hetoroporous heterogeneous ceramics showed high radiations shielding capabilities up to 2000 °C in vacuum. Even after repeated exposures at higher temperatures, a significant degradation of the SiSiC scaffold was not observed.

To improve the thermo-chemical resistance of PIP–C f /SiC composites, the SiC matrix is modified by adding ZrB 2 and Ta powder to the pre-ceramic slurry to form C f /SiC–ZrB 2 –TaC composites. Within this study the modified composites are investigated regarding their microstructure, chemical composition and physical properties (density = 2,39–2,72 g/cm 3 ; porosity = 20,3–24,8 vol.-%; fiber volume content = 52–57 vol.-%). Mechanical properties are investigated in order to ensure that there is no negative influence by ZrB 2 and TaC matrix modification. The matrix modification is followed by an improvement in bending strength (up to 27% increase), Young’s modulus (up to 28% increase) and for interlaminar shear strength (up to 22% increase). Finally the thermo-chemical behavior of the C f /SiC–ZrB 2 –TaC composites is evaluated in a combustion chamber-like environment using the Airbus Group long-term material test facility (Environmental Relevant Burner Rig-Kerosene, ERBURIG K ). The results show that the thermo-chemical resistance of C f /SiC–ZrB 2 –TaC composites is improved and the oxygen permeability through the composite is decreased (from 5 to 1 layer).

Industrially produced Cf/C ceramic composites have been brazed to Nimonic alloys using a TiCuSil filler metal. Ιn order to accommodate the different linear coefficients of expansion between ceramic and metal as well as to provide compatibility between the surfaces to be joined, the Cf/C surface was metallized through the deposition of a chromium layer. Subsequent heat treatments were carried out to develop intermediate layers of chromium carbides. Crack-free joints have been produced and shear tests show that failure occurs within the composite. At the Cf/C-filler interface a layered structure of the metallic elements is observed. Titanium is depleted from the filler zone and interacts with the carbon to form carbides. In the filler region, Ag and Cu rich regions are formed.

Hypersonic airliner would be exposed to temperatures that are beyond the limits of classical aircraft materials. In order to handle this problem the latest developments of new materials and composite structures suitable for high temperature application need to be taken into account. The focus of the European Research program ATLLAS is on advanced light-weight, high-temperature material development strongly linked to a high-speed passenger aircraft design. ATLLAS stands for Aerodynamic and Thermal Load Interactions with Lightweight Advanced Materials for High Speed Flight. The 4 years program ATLLAS-II is a logical continuation project built upon the experience and technology development gained within ATLLAS-I. The corresponding work related to combustor structures and material integration deals with the opportunity to investigate at academic level, both in basic and relevant environment, different solutions possibly usable to ensure the long range cruise of a high speed airliner. Different materials (UTHC, CMC, metallic) and different cooling techniques (radiation, convective, transpiration) are studied. Available numerical or semi-empirical tools are used to prepare the test, to design the different architectures. A pin fin experiment allows to better know the pressure drop and the heat transfer for different channel patterns with thermal cristal techniques. The ERBURIGK long duration test facility allows to characterize different ceramic matrix composite uncooled samples and will allow to realize, at small scale, a long duration (several hours) investigation of cooled ceramic structure in PTAH-SOCAR technology. A multifunctional metallic transpiration cooled HSS panel using Hollow Spheres Stacking as core material was designed and preliminary tested in cold conditions with GN2 and in hot conditions with infra red lamps under 1 MW/m² heat flux. CMC and UHTC materials are used to design, manufacture and test generic fin injectors usable in high speed combustors. Industrial hypersonic METHYLE test facility is used to test in relevant Mach 6 combustor environment HSS panel as well as advanced fin injectors. Hot and cold permeability of composites is documented with GN2 and GH2. Numerical models are used in accordance with the experiments, some examples are given in the present paper.

Mechanical performance of three oxide/oxide ceramic matrix composites (CMCs) based on Nextel 610 fibers and SiOC, alumina, and mullite/SiOC matrices respectively, is evaluated herein. Tensile strength and stiffness of all materials decreased at 1000 °C and 1200 °C, probably because of degradation of fiber properties beyond 1000 °C. Microstructural changes in the composites during exposure at 1000 °C and 1200 °C for 50 h reduce their flexural strength, fracture toughness and work of fracture. A literature review regarding mechanical properties of several oxide/oxide CMCs revealed lower influence of fiber properties on composite strength compared with elastic modulus. The tested composites exhibit comparable stiffness and strength but higher fracture toughness compared with average values determined from a literature review. Considering CMCs with different compositions, we observed an interesting linear trend between strength and fracture toughness. The validity of the linear relationship between fracture strength and flexural toughness for CMCs is discussed.

Blanching of copper alloys is a major life-limiting degradation mechanism in rocket combustion chambers. Understanding of this mechanism is key to further improvement of rocket engine performance, so to gain further insight into the mechanism oxidationreduction cycles with a variation of temperature, amount of cycles and reduction atmosphere composition have been used to duplicate the blanching mechanism on sample scale. As part of the evaluation the sample weight and the roughness have been measured in the initial state as well as after testing. Furthermore, scanning electron microscope (SEM) investigations have been conducted to distinguish whether sponge-like blanching structures formed during testing. The roughness measurements showed a smoothing of the surface up to about 600°C, with a maximum at about 500°C. Above these temperatures the roughness increased rapidly. The SEM investigation proved to deliver sponge-like blanching structures, also with a strong dependence on temperature. The experiments proved to give trends on the influence of the individual parameters. Temperature was found to be the principal influencing factor, followed by the duration (amount of oxidation-reduction cycles). The composition of the reduction atmosphere was found to play only a supporting role. The results obtained in this study will be the baseline for further investigations in the newly developed Airbus Group Innovations ERBURIG H test facility.

Hypersonic airliner would be exposed to temperatures that are beyond the limits of classical aircraft materials. In order to handle this problem the latest developments of new materials and composite structures suitable for high temperature application need to be taken into account. The focus of the European Research program ATLLAS is on advanced light-weight, high-temperature material development strongly linked to a high-speed passenger aircraft design. ATLLAS stands for Aerodynamic and Thermal Load Interactions with Lightweight Advanced Materials for High Speed Flight. The 4 years program ATLLAS-II is a logical continuation project built upon the experience and technology development gained within ATLLAS-I. The corresponding work related to combustor structures and material integration deals with the opportunity to investigate at academic level, both in basic and relevant environment, different solutions possibly usable to ensure the long range cruise of a high speed airliner. Different materials (UTHC, CMC, metallic) and different cooling techniques (radiation, convective, transpiration) are studied. Available numerical or semi-empirical tools are used to prepare the test, to design the different architectures. A pin fin experiment allows to better know the pressure drop and the heat transfer for different channel patterns with thermal cristal techniques. The ERBURIG-K long duration test facility allow to characterize different ceramic matrix composite uncooled samples and will allow to realize, at small scale, a long duration (several hours) investigation of cooled ceramic structure in PTAH-SOCAR technology. A multifonctionnal metallic transpiration cooled HSS panel using Hollow Spheres Stacking as core material was designed and preliminary tested in cold conditions with GN2 and in hot conditions with infra lamps under 1 MW/m² heat flux. CMC and UHTC materials are used to design, manufacture and test generic fin injectors usable in high speed combustors. Industrial hypersonic METHYLE test facility is used to test in relevant Mach 6 combustor environment HSS panel as well as advanced fin injectors. Hot and cold permeability of composites is documented with GN2 and GH2. Numerical models are used in accordance with the experiments, some examples are given in the present paper.

This study analyses the effects of heat treatments in oxidative atmosphere on the mechanical and microstructural properties of a fiber-reinforced weak interface composite (UMOX™) which is composed of a mullite-SiOC matrix and Nextel™ 610 fibers with fugitive coatings. Composites of different porosity grades, depending on the polymer infiltration and pyrolysis cycle, are exposed to 1000 and 1200 °C for 50 h. The exposure provokes the formation of silica, which leads to matrix densification and the formation of silica bridges at the fiber–matrix interface, resulting in an increased interfacial bonding strength. Consequently, the fracture toughness and the flexural strength are significantly reduced. The study confirms that SiOC-based materials are suitable for an application at high temperatures in oxygen-rich atmospheres up to 1000 °C. It is, however, important to consider the microstructural changes and thereby induced decrease of the overall mechanical performance during a high-temperature use.

The tensile and creep performance of two Nextel® 610 (N610) reinforced polymer derived ceramic composites (UMOX™, OXIPOL) is studied up to 1200 °C. Independent of the fiber orientation (±45° or 0°/90°) all samples exhibit a segment where the strain rate was constant. The creep performance in ±45° is matrix dominated and shows a more pronounced primary creep regime, due to changes within the matrix. The following creep regime with a constant strain rate might be attributed to viscous flow of the SiOC within the matrix based on activation energy (283 kJ/mol) and stress exponent (0.6). In 0°/90° orientation the creep and tensile performance is independent of oxidation, but directly influenced by grain structure of the fiber. The coarser and non-uniform microstructure of the fibers in UMOX™ decreases the stationary creep rates and changes the diffusional creep mechanism. The possibility to modify the microstructure of the fiber during the manufacturing process might be used to adjust e.g. the strength and creep stability of these materials related to the desired applications.

In the framework of the High Performance Oxide Ceramics program (HiPOC), three different oxide/oxide ceramic matrix composite (CMC) materials are studied for a combustion chamber application in continuation of the work reported in Gerendas et al. [1]. A variation in the micro-structural design of the three CMC materials in terms of different fiber architecture and matrix processing are considered in a first work stream. By modification of the matrix and the fiber-matrix interface as well as the application of an environmental barrier coating (EBC), the high temperature stability is enhanced. Furthermore, design concepts for the attachment of the CMC component to the metal structure of the engine are finalized in a second work stream. Issues like sealing of cooling leakage paths, allowance for the different thermal expansion and the mechanical fixation are addressed. An interim standard of the mechanical attachment scheme is studied on a shaker table. Also the friction coefficient between the metallic and ceramic components is analyzed in order to set the proper tightening torque. The manufacturing of the CMC combustor is improved in several iterations in order to achieve a high quality material with optimized fiber architecture. Afterwards, two CMC materials are selected for the combustion testing and the finalized design of the metallic and CMC components is manufactured. A fit check is performed prior to EBC application and laser drilling of the effusion holes in order to evaluate the impact of the manufacturing tolerances on the function of the sealing and attachment scheme and to correct small issues at this stage. First results from the validation testing in a high-pressure tubular combustion rig up to a Technology Readiness Level 4 (TRL4) are reported.Copyright © 2013 by ASME and Rolls-Royce Deutschland Ltd & Co KG

Combustion chambers of orbital thrusters and components of dual-mode ramjets require materials which are able to withstand high aerodynamic and thermo-mechanical loads during typical operation. Higher combustion efficiency and/or performance lead to higher combustion temperatures and therefore higher material working temperatures in the range of 1600°C to 1900°C. In addition to that severe thermo-chemical interactions between the combustion products and the substrate (e.g. oxidation, carburetion, nitruration) complete this harsh environment. Such demanding requirements in combination with a high thrust to weight ration limit the field of possible material candidates to ultra-high-temperature ceramics, ceramic matrix composites (CMCs) (e.g. C/SiC, C/C-SiC...) and coating systems (e.g. β-SiC, HfC, ZrB2, HfB2...). Using the novel established EADS ERBURIG K

Three different oxide/oxide ceramic matrix composite (CMC) materials are described. Design concepts for the attachment of the CMC component to the metal structure of the gas turbine are developed in a first work stream focused on the combustion chamber and the turbine seal segment. Issues like environmental barrier coating (EBC)/thermal barrier coatings (TBC), application and volatilization, allowance for the different thermal expansion and the mechanical fixation are addressed. The design work is accompanied by CFD and FEM simulations. A variation of the microstructural design of the three oxide/oxide CMC materials in terms of different fiber architecture and processing of matrix are considered. Also, mechanical properties of these variations are evaluated. The material concepts are developed further in a second work stream. The CMCs are tested in various loading modes (tension, compression, shear, off-axis loading) from room temperature to maximum application temperature focusing on tensile creep behavior. By modification of the matrix and the fiber-matrix interface as well as EBC coatings, the high temperature stability and the insulation performance are enhanced. An outline of the “High Performance Oxide Ceramic”-program HiPOC for the following years is given, including manufacturing of a high-pressure tubular combustor and turbine seal segments from the improved materials as technology samples, for which validation testing up to technology readiness level 4 is scheduled for 2011.Copyright © 2011 by ASME and Rolls-Royce Deutschland Ltd & Co KG

Barley beta-glucan solubilase is an enzyme that degrades barley beta-glucan in extracts obtained from barley flour. The solubilase preferentially attacks the longer blocks of beta-(1-->4) linkages, i.e., those containing at least nine glucosyl residues. There is strong evidence to suggest that the solubilase derives from fungi associated with the husk of the grain. It was found that cellulase (EC 3.2.1.4) from Trichoderma sp. shows similar activity under dilute conditions. Since fungi associated with the husk of the grain are known to produce these types of cellulases, there is no need, based on current evidence, to propose the existence of a unique enzyme, i.e., solubilase, for the solubilising behaviour of enzymes in the barley grain.