The purpose of this paper is to present a multidisciplinary predesign process and its application to three aero-engine models. First, a 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 (CROR) and an ultrahigh 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 predesign 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 different fidelity levels applied within the design phases. To cope with the inherent complexity, data modeling techniques have been applied to explicitly determine 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 data model, the entire engine predesign process is presented. Starting with the definition of a flight mission scenario and resulting top level engine requirements, thermodynamic engine performance models are developed. By means of these thermodynamic models, a detailed engine component predesign 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 context of the overall engine system.

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Over the past decades of preliminary aero-engine design a great effort has been invested in increasing steady-state efficiency to reduce missions fuel burn and thus CO2 emissions. Whilst pushing the performance cycle further to its limits previously minor deemed processes such as the transient behavior become more important because engines stability must still be granted for. It is therefore beneficial to integrate transient performance early into the overall aero-engine design tool chain to be able to predict the dynamic behavior from the beginning. This paper presents a methodology to couple predesign and transient performance simulation in order to get a more holistic picture of aero-engines early in the design process. This procedure entails extensive amount of data transfer throughout multi-disciplinary tools with different fidelity levels. This task is tackled using DLRs virtual engine platform GTlab (Gas Turbine Laboratory), which provides a geometric data model with abstract description of predesign components and standardized interfaces for data exchange. In order to demonstrate the proposed methodology a performance model of a turbofan similar to the V2500 aero-engine is used. For that purpose, a performance cycle is established providing boundary conditions for the preliminary aerodynamic engine design. The designed components provide necessary input data for the subsequent transient certification maneuver Eventually, parametric studies are conducted to show the impact of design variations on transient data such as the minimum surge margin and minimum tip clearance as well as on preliminary engine design.

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Due to a high degree of complexity and computational effort, overall system simulations of jet engines are typically performed as 0-dimensional thermodynamic performance analysis. Within these simulations and especially in the early cycle design phase, the usage of generic component characteristics is common practice. Of course these characteristics often cannot account for true engine component geometries and operating characteristics which may cause serious deviations between simulated and actual component and overall system performance. This leads to the approach of multi-fidelity simulation, often referred to as zooming, where single components of the thermodynamic cycle model are replaced by higher-order procedures. Hereby the consideration of actual component geometries and performance in an overall system context is enabled and global optimization goals may be considered in the engine design process. The purpose of this study is to present a fully automated approach for the integration of a 3D-CFD component simulation into a thermodynamic overall system simulation. As a use case, a 0D-performance model of the IAE-V2527 engine is combined with a CFD model of the appropriate fan component. The methodology is based on the DLR in-house performance synthesis and preliminary design environment GTlab combined with the DLR in-house CFD solver TRACE. Both, the performance calculation as well as the CFD simulation are part of a fully automated process chain within the GTlab environment. The exchange of boundary conditions between the different fidelity levels is accomplished by operating both simulation procedures on a central data model which is one of the essential parts of GTlab. Furthermore iteration management, progress monitoring as well as error handling are part of the GTlab process control environment. Based on the CFD results comprising fan efficiency, pressure ratio and mass flow, a map scaling methodology as it is commonly used for engine condition monitoring purposes is applied within the performance simulation. Hereby the operating behavior of the CFD fan model can be easily transferred into the overall system simulation which consequently leads to a divergent operating characteristic of the fan module. For this reason, all other engine components will see a shift in their operating conditions even in case of otherwise constant boundary conditions. The described simulation procedure is carried out for characteristic operating conditions of the engine.

For an efficient detection of single or multiple component damages, the knowledge of their impact on the overall engine performance is crucial. This knowledge can be either built up on measurement data, which is hardly available to non-manufacturers or –maintenance companies, or simulative approaches such as high fidelity component simulation combined with an overall cycle analysis. Due to a high degree of complexity and computational effort, overall system simulations of jet engines are typically performed as 0-dimensional thermodynamic performance analysis, based on scaled generic component maps. The approach of multi-fidelity simulation, allows the replacement of single components within the thermodynamic cycle model by higher-order simulations. Hence, the component behavior becomes directly linked to the actual hardware state of the component model. Hereby the assessment of component deteriorations in an overall system context is enabled and the resulting impact on the overall system can be quantified. The purpose of this study is to demonstrate the capabilities of multi fidelity simulation in the context of engine condition monitoring. For this purpose, a 0D-performance model of the IAE-V2527 engine is combined with a CFD model of the appropriate fan component. The CFD model comprises the rotor as well as the outlet guide vane of the bypass and the inlet guide vane of the core section. As an exemplarily component deterioration, the fan blade tip clearance is increased in multiple steps and the impact on the overall engine performance is assessed for typical engine operating conditions. The harmonization between both simulation levels is achieved by means of an improved map scaling approach using an optimization strategy leading to practicable simulation times.

Variable geometry blade rows are a common instrument to avoid compressor instabilities which occur especially at low- and full-speed operation of gas turbines. The operating settings of variable stator vanes (VSVs) are typically obtained from expensive and time consuming performance rig tests and are not known during the early design phase of a gas turbine. During preliminary design of the overall engine it is common practice to use default component characteristics based on considerable engineering experience. These can deviate substantially at off-design and often do not properly account for the impact of changes in component geometry. As a solution, multi-fidelity simulation often referred to as zooming or variable complexity analysis is applied. This proceeding facilitates a transfer of single component performance characteristics obtained in mid- or high-fidelity analysis to a full gas turbine system analysis based on lower resolution level. The purpose of this study is to present a multidisciplinary numerical optimization methodology to define ideal blade row staggering of variable compressor stator vanes during the early preliminary design phase using multi-fidelity simulation. The objective of the resultant multi-dimensional constraint optimization is to find the best solution for the entire gas turbine system for a set of discrete operating points. For the assessment a generic turbofan engine model is designed by taking into account top level engine requirements from an assumed airframe and flight mission scenario. Based on the performance calculation a full 3-D axial multistage high pressure compressor (HPC) is designed. The assumed design considerations are summarized and the modelling techniques are presented. The optimization of VSV staggering mentioned above is carried out by re-staggering the variable geometry blade rows of the high-fidelity HPC and run a full 2-dimensional through-flow calculation. Results are then automatically transferred to the 0-dimensional engine model to calculate the engine overall performance. A Pareto optimized blade row staggering is found by taking into account the surge margin and the specific fuel consumption of the entire engine system as objective functions of the optimization process. Simultaneously several constraints such as DeHaller numbers and diffusion factors are considered. The optimization process chain and the tool coupling are summarized and described in detail. The resulting VSV staggering for a set of discrete operating points is shown.

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