Design Optimization of a Multi-Stage Axial Compressor Using Throughflow and a Database of Optimal Airfoils

The basic tool set to design multi-stage axial compressors consists of fast codes for throughflow and blade-to-blade analysis. Detailed blade row design is conducted with 3D CFD, mainly to control the end wall flow. This work focuses on the interaction between throughflow and blade-to-blade design and the transition to 3D CFD. A design strategy is presented that is based on a versatile airfoil family. The new class of airfoils is generated by optimizing a large number of airfoil shapes for varying design requirements. Each airfoil geometry satisfies the need for a wide working range as well as low losses. Based on this data, machine learning is applied to estimate optimal airfoil shape and performance. The performance prediction is incorporated into the throughflow code. Based on a throughflow design, the airfoils can be stacked automatically to generate 3D blades. On this basis, a 3D CFD setup can be derived. This strategy is applied to study upgrade options for a 15-stage stationary gas turbine compressor test rig. At first, the behavior of the new airfoils is studied in detail. Afterwards, the design is optimized for mass flow rate as well as efficiency. Selected configurations from the Pareto-front are evaluated with 3D CFD

Influence of Different Flow Solvers and Off-Design Conditions on the Determination of Fan-Rotor Wakes for Broadband Noise Prediction

The acoustic interaction of fan-rotor wakes with the downstream stator vanes is considered as an important noise source of an aircraft engine. The turbulence induced by the rotor generates a stochastic acoustic source that appears as broadband noise in the acoustic spectrum. During the preliminary design phase of an engine, established meanline and throughflow solvers usually do not resolve turbulence and associated unsteady flow parameters. But such solvers provide rotor pressure losses that can be used to estimate the mean and turbulent rotor wakes. A crucial step is the deduction of turbulence parameters from the mean wakes. A semi-empirical model for rotor-wake turbulence estimation is presented in this paper. The meanline method and the throughflow solver are compared to three-dimensional computational flow simulations investigating the capabilities of the different solvers to provide flow data for broadband wake interaction noise prediction. The methods are applied to a representative modern fan stage at a comprehensive number of operating points, comprising several speed lines from surge to choking conditions. Microphone measurements are consulted to assess the noise predictions. The evaluation confirms the applicability of the meanline and throughflow method in combination with the turbulence model for broadband noise estimation during the preliminary design phase. The underestimated turbulence in the tip region of the fan is found to be negligible even during off-design conditions

Exploring a Database of Optimal Airfoils for Axial Compressor Design

Ensuring a high degree of commonality among a range of products can dramatically decrease development costs. This paper aims to generate a highly versatile compressor airfoil family that covers most applications in the core compression system of aircraft engines and stationary gas turbines. This airfoil family is generated by filling a database with optimized airfoil shapes. The database is structured in seven dimensions, denominated as 'design requirements': blade stagger angle, pitch-chord ratio, profile area and the following design point properties: inlet Mach number, Reynolds number, streamtube contraction and aerodynamic loading. Additional constraints are imposed to ensure that feasible airfoils exist for each set of requirements. These constraints include limitations for profile area depending on inlet Mach number and limits for axial Mach number. To fill this seven-dimensional space, a large number of airfoils is generated by means of numerical optimization at discrete points in this space. The target is to find airfoil shapes that have low losses and ensure stable operation over wide incidence ranges. Design and off-design performance is evaluated with the blade-to-blade flow solver MISES. The solver is well established among industry and research and it is validated to a high degree by experiments. To verify the optimization strategy, it is tested on a set of existing compressor airfoils. The optimized geometries of four of the airfoils under investigation are found in the appendix. The database offers a wide variety of airfoils for different applications. Airfoils for sub- and supersonic inflow are covered as well as airfoils suited for placement at hub or casing. The benefit of using airfoils optimized for their specific purpose over having generic airfoil shapes is discussed as well. In future, this airfoil database will be used to study novel concepts for aircraft engines

Automated Calibration of Compressor Loss and Deviation Correlations

Throughout preliminary design of multistage axial compressors performance prediction is based on fast 2D throughflow codes incorporating models for loss, deviation and secondary flow effects. To achieve high solution quality these models have to be calibrated on experimental or simulation data. Particularly 2D cascade loss and deviation correlations are of major importance. Correlations available in literature are mostly based on data from early cascade wind tunnel tests of NACA-65 and DCA profile families. In general, calibrations for state-of-the-art airfoil shapes are not published. This paper describes a methodology to re-calibrate parameters of well-known empirical loss and deviation correlations to CFD simulations of compressor cascades. Correlations describing choke incidence, on- and off-design loss and deviation are presented, resolving dependencies on inlet flow angle, Mach number, Reynolds number and MVDR. Based on a set of loss and deviation polars, nonlinear least-squares problems are solved to determine the empirical parameters. By including shock losses and choking incidence, the model is suited for subsonic as well as supersonic inlet conditions. Results of the calibration process are shown for two state-of-the-art compressor cascades. The overall model is evaluated on a sub- and a transonic compressor as well as on a transonic fan comparing throughflow results with 3D CFD and experimental data. A high level of automation makes the methodology applicable to arbitrary profile families or in combination with programs for parametric airfoil design

A Database of Optimal Airfoils for Axial Compressor Throughflow Design

This text describes methods to organize a large set of optimized airfoils in a relational database and its application in throughflow design. Optimized airfoils are structured in five dimensions: inlet Mach number, blade stagger angle, pitch-chord ratio, maximum thickness-chord ratio and a parameter for aerodynamic loading. In this space, a high number of airfoil geometries is generated by means of numerical optimization. Each airfoil geometry is tailored to its specific requirements and optimized for a wide working range as well as low losses. During the optimization of each airfoil, performance in design and off-design conditions is evaluated with the blade-to-blade flow solver MISES. Together with airfoil geometry, the database stores automatically calibrated correlations which describe cascade performance in throughflow calculation. Based on these methods, two subsonic stages of a 4.5-stage transonic research compressor are redesigned. Performance of baseline and updated geometries is evaluated with 3D CFD. The overall approach offers accurate throughflow design incorporating optimized airfoil shapes and a fast transition from throughflow to 3D CFD design

Design optimization of a multi-stage axial compressor using through flow and a database of optimal airflows

The basic tool set to design multi-stage axial compressors consists of fast codes for throughflow and blade-to-blade analysis. Detailed blade row design is conducted with 3D CFD, mainly to control the end wall flow. This work focuses on the interaction between throughflow and blade-to-blade design and the transition to 3D CFD. A design strategy is presented that is based on a versatile airfoil family. The new class of airfoils is generated by optimizing a large number of airfoil shapes for varying design requirements. Each airfoil geometry satisfies the need for a wide working range as well as low losses. Based on this data, machine learning is applied to estimate optimal airfoil shape and performance. The performance prediction is incorporated into the throughflow code. Based on a throughflow design, the airfoils can be stacked automatically to generate 3D blades. On this basis, a 3D CFD setup can be derived. This strategy is applied to study upgrade options for a 15-stage stationary gas turbine compressor test rig. At first, the behavior of the new airfoils is studied in detail. Afterwards, the design is optimized for mass flow rate as well as efficiency. Selected configurations from the Pareto-front are evaluated with 3D CFD

Influence of Different Flow Solvers and Off-Design Conditions On the Determination of Fan-Rotor Wakes for Broadband Noise Prediction

The acoustic interaction of fan-rotor wakes with the downstream stator vanes is considered as an important noise source of an aircraft engine. The turbulence induced by the rotor generates a stochastic acoustic source that appears as broadband noise in the acoustic spectrum. During the preliminary design phase of an engine, established meanline and throughflow solvers usually do not resolve turbulence and associated unsteady flow parameters. But such solvers provide rotor pressure losses that can be used to estimate the mean and turbulent rotor wakes. A crucial step is the deduction of turbulence parameters from the mean wakes. A semi-empirical model for rotor-wake turbulence estimation is presented in this paper. The meanline method and the throughflow solver are compared to three-dimensional computational flow simulations investigating the capabilities of the different solvers to provide flow data for broadband wake interaction noise prediction. The methods are applied to a representative modern fan stage at a comprehensive number of operating points, comprising several speed lines from surge to choking conditions. Microphone measurements are consulted to assess the noise predictions. The evaluation confirms the applicability of the meanline and throughflow method in combination with the turbulence model for broadband noise estimation during the preliminary design phase. The underestimated turbulence in the tip region of the fan is found to be negligible even during off-design conditions

Design of the Compression System of a Geared Turbofan

The geared turbofan aircraft engine is an efficient type of aircraft engine. By using a gearbox, the fan can run at a different rotational speed in comparison to the shaft connecting low pressure compressor and turbine. To assess the in-house tools for preliminary design, a geared turbofan, with the same requirements as the Pratt&Whitney 1000G, is designed, including thermodynamic cycle, turbine, compressor, burner and structural mechanics. The goal of this study is the design of the compression system, composed of the low-pressure compressor, the inter-compressor duct and the high-pressure compressor. The following steps have been carried out to design the compression system: first, a preliminary design of both compressors is realized, following by a detailed design, conducted with throughflow calculations and the use of a database of optimal airfoils [1]. Three stages have been chosen for the low-pressure compressor and eight stages for the high-pressure compressor. In order to reach the target conditions for all operating points from the specifications, the schedule for the variable guide vanes has also been decided on. The final design is validated with 3D CFD calculations. Additionally, the interaction of the inter-compressor duct with the low-pressure and the highpressure compressor is investigated

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