Richard-Gregor Becker

An Integrated Method for Propulsion System Conceptual Design

The conceptual design of future, potentially highly integrated aircraft engines pose a variety of new design options to the propulsion system engineers. In order to find the best conceptual design, rapid evaluation of many design choices is essential. However, traditional, fast evaluation methods employing historical and empirical data can only be applied to novel engine concepts to a very limited degree. Thus, swift conceptual design methods based on physical approaches providing a sophisticated level of detail are needed. The current paper presents a methodology focused on conceptual engine design. The methodology is based on the gas turbine simulation framework GTlab, which integrates software tools for engine performance, component aerodynamics and structural design. For conceptual design a dedicated set of design tools exists — the so called GTlab-Sketchpad. Sketchpad tools have full access to the thermodynamic design data of the engine performance module. Based on cycle analysis, the tool set generates parametric representations of the propulsion system components and stores the results back to the frameworks data model. Computational time is limited to a few seconds, to ensure interactivity during the design process. The graphical user interface provides means to interactively modify the design parameters and to immediately evaluate their impact on the overall design. Since the internal data model facilitates three dimensional parameterizations of the engine components, 3D representations of the engine designs can be generated by interfacing an open source CAD-kernel. For the present paper, the conceptual design process of a commercial jet engine utilizing GTlab-Sketchpad is shown. The underlying computational methods are described and the resulting 3D-geometry is presented.Copyright

Multi-Level MDO of a Long-Range Transport Aircraft Using a Distributed Analysis Framework

DLRs work on developing a distributed collaborative MDO environment is presented. A multi-level Approach combining high-fidelity MDA for aerodynamics and structures with conceptual aircraft design methods is employed. Configuration-specific sizing loads are evaluated and used for sizing the structure. A gradient-free optimization algorithm is used to optimize the fuel burn of a generic long-range wide-body transport aircraft configuration with 9 shape parameters. The results show a truly multidisciplinary improvement of the modified design. The result of a gradient-free high-fidelity MDO with preselected load cases and five shape parameters is also presented, comparing a full mission analysis with results for the Breguet range equation.

Application of Kriging Metamodels to the Automated Start Value Generation for Gas Turbine Performance Simulations

This paper presents a process for automated start value generation for gas turbine performance simulations using Kriging metamodels. The metamodels are trained on a small number of pre-selected operating points in the flight envelope. Predictions of the trained metamodels are used as initialization parameters for subsequent performance simulations of arbitrary operating points in order to increase robustness and computational speed of the numerical process. Different approaches for the selection of the training points are evaluated. A comparison to the classical approach of tablebased initialization is carried out to highlight the advantages and disadvantages of the new methodology. Furthermore, the inclusion of supplementary operating points into the training sample of the metamodels is analyzed. Depending on certain criteria, such as the difference of prediction and simulation result, operating points are included into the sample and a retraining of the metamodels is performed. Simulations using the retrained models as guess value generators are compared to the previous Kriging approach. The advantages and disadvantages of the retraining approach are discussed.

Reproducing Existing Nacelle Geometries With the Free-Form Deformation Parametrization

In the conceptual and preliminary aircraft design phase the Free-Form Deformation (FFD) is one of various parametrization schemes to define the geometry of an engines nacelle. In this paper we present a method that is able to create a G2 continuous approximation of existing reference nacelles with the B-spline based FFD, which is a generalization of the classical FFD. The basic principle of our method is to start with a rotational symmetric B-spline approximation of the reference nacelle, which is subsequently deformed with a FFD grid that is placed around the initial geometry. We derive a method that computes the displacement of the FFD grid points, such that the deformed nacelle approximates the reference nacelle with minimal deviations. As this turns out to be a linear inverse problem, it can be solved with a linear least squares fit. To avoid overfitting effects - like degenerative FFD grids which imply excessive local deformations - we regularize the inverse problem with the Tikhonov approach. For the reference geometries we selected the NASA CRM model and the IAE V2500 engine. Both resemble nacelles that are typically found on common aircraft models and both deviate sufficiently from the rotational symmetry. We demonstrate that the mean error of our approximation decreases with an increase of the number of FFD grid points and how the regularization affects these results. Finally, we compare the B-spline based FFD with the classical Bernstein based FFD for both models.

COMPARISON OF A HEAT SOAKAGE MODEL WITH TURBOFAN TRANSIENT ENGINE DATA

In the present paper, a transient performance code is employed to predict on-wing test data of the IAE-V2500 engine mounted on an Airbus A320-232. The test data was recorded by the engine control system and may serve as an open basis for validation of future transient studies. For the current investigation, the employed code considers the fundamental equations of the constant mass flow method as well as heat transfer effects by a lumped parameter approach. The study focuses on seven accelerations and one deceleration. Engine test data was gathered with 10Hz sampling rate, imprinting the applied time step of the model. First, the steady-state matching of the test data was conducted. Subsequently, the measurement quantities fuel flow, inlet temperature and inlet pressure were prescribed as time-varying boundary conditions to the transient model. The results of the standard transient model and the model including thermal effects were compared with temperatures, pressures and shaft speeds. The LPT outlet temperature and the working line excursion in the booster map were examined in detail. The outcome concurs with the original statement that thermal effects are mandatory to enhance model accuracy. Lastly, a sensitivity analysis of the thermal input parameters was accomplished and its influence on model prediction investigated.

Comparison of Breguet and ODE Evaluation of the Cruise Mission Segment in the Context of High-Fidelity Aircraft MDO

This report presents a multi-disciplinary optimization (MDO) process that minimizes the mission fuel burn of an aeroelastic long-range transport aircraft configuration, by modifying the wing planform, twist, and structural element thicknesses. Two optimizations are performed, one where the fuel burn is approximately evaluated through Breguet range equation, and the other where the ordinary differential equation (ODE) for the step-climb cruise is formally integrated. This is done in order to determine if the Breguet equation is still sufficient in face of high-fidelity aeroelastic simulations. The two optimized designs ended up having similar improvements, thus confirming the applicability of the Breguet equation, for the number of design parameters that were employed.

OPTIMIZATION OF COMPRESSOR VARIABLE GEOMETRY SETTINGS USING MULTI-FIDELITY SIMULATION

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.