Sandrine Hönig

Development and Test of Oxide/Oxide Ceramic Matrix Composites Combustor Liner Demonstrators for Aero-engines

Ceramic matrix composites (CMC) offer the potential of increased service temperatures and are thus an interesting alternative to conventional combustor alloys. Tubular combustor liner demonstrators made of an oxide/oxide CMC were developed for a lean combustor in a future aero-engine in the medium thrust range and tested at engine conditions. During the design various aspects like protective coating, thermo-mechanical design, development of a failure model for the CMC as well as design and test of an attachment system were taken into account. The tests of the two liners were conducted at conditions up to 80% take-off. A new protective coating was tested successfully with a coating thickness of up to t=1 mm. Different inspection criteria were derived in order to detect crack initiation at an early stage for a validation of the failure model. With the help of detailed pre- and post-test computer tomography scans to account for the micro structure of the CMC the findings of the failure model were in reasonable agreement with the test results.

DEVELOPMENT AND TEST OF OXIDE/OXIDE CMC COMBUSTOR LINER DEMONSTRATORS FOR AERO ENGINES

Ceramic matrix composites (CMC) offer the potential of increased service temperatures and are thus an interesting alternative to conventional combustor alloys. Tubular combustor liner demonstrators made of an oxide/oxide CMC were developed for a lean combustor in a future aero-engine in the medium thrust range and tested at engine conditions. During the design various aspects like protective coating, thermo-mechanical design, development of a failure model for the CMC as well as design and test of an attachment system were taken into account. The tests of the two liners were conducted at conditions up to 80% take-off. A new protective coating was tested successfully with a coating thickness of up to t=1 mm. Different inspection criteria were derived in order to detect crack initiation at an early stage for a validation of the failure model. With the help of detailed pre- and post-test computer tomography scans to account for the micro structure of the CMC the findings of the failure model were in reasonable agreement with the test results.

DESIGN AND APPLICATION OF A MULTI-DISCIPLINARY PRE-DESIGN PROCESS FOR NOVEL ENGINE CONCEPTS

Central targets for jet engine research activities comprise the evaluation of improved engine components and the assessment of novel engine concepts for enhanced overall engine performance in order to reduce the fuel consumption and emissions of future aircraft. Since CO2 emissions are directly related to engine fuel burn, a reduction in fuel consumption leads to lower CO2 emissions. Therefore improvements in engine technologies are still significant and a multi-disciplinary pre-design approach is essential in order to address all requirements and constraints associated with different engine concepts. Furthermore, an increase in effectiveness of the preliminary design process helps reduce the immense costs of the overall engine development. Within the DLR project PEGASUS (Preliminary Gas Turbine Assessment and Sizing) a multi-disciplinary pre-design and assessment competence of the DLR regarding aero engines and gas turbines was established. The application of modern preliminary design methods allows for the construction and evaluation of innovative next generation engine concepts. The purpose of this paper is to present the developed multi-disciplinary pre-design process and its application to three aero engine models. First, a state of the art twin spool mixed flow turbofan engine model is created for validation purposes. The second and third engine models investigated comprise future engine concepts: a Counter Rotating Open Rotor and an Ultra High Bypass Turbofan. The turbofan used for validation is based on publicly available reference data from manufacturing and emission certification. At first the identified interfaces and constraints of the entire pre-design process are presented. An important factor of complexity in this highly iterative procedure is the intricate data flow, as well as the extensive amount of data transferred between all involved disciplines and among the different fidelity levels applied within the smoothly connected design phases. To cope with the inherent complexity data modeling techniques have been applied to explicitly determine the required data structures of those complex systems. The resulting data model characterizing the components of a gas turbine and their relationships in the design process is presented in detail. Based on the established data model the entire engine pre-design process is presented. Starting with the definition of a flight mission scenario and the resulting top level engine requirements thermodynamic engine performance models are developed. By means of these thermodynamic models, a detailed engine component pre-design is conducted. The aerodynamic and structural design of the engine components are executed using a stepwise increase in level of detail and are continuously evaluated in the context of the overall engine system.