Heinz-Adolf Schreiber

1999 Turbomachinery Committee Best Paper Award: Development of Advanced Compressor Airfoils for Heavy-Duty Gas Turbines— Part I: Design and Optimization

A new family of subsonic compressor airfoils, which are characterized by low losses and wide operating ranges, has been designed for use in heavy-duty gas turbines. In particular the influence of the higher airfoil Reynolds numbers compared to aeroengine compressors and the impact of these differences on the location of transition are taken into account. The design process itself is carried out by the combination of a geometric code for the airfoil description, with a blade-to-blade solver and a numerical optimization algorithm. The optimization process includes the design-point losses for a specified Q3D flow problem and the off-design performance for the entire operating range. The family covers a wide range of inlet flow angle, Mach number, flow turning, blade thickness, solidity and AVDR in order to consider the entire range of flow conditions that occur in practical compressor design. The superior performance of the new airfoil family is demonstrated by a comparison with conventional controlled diffusion airfoils (CDA). The advantage in performance has been confirmed by detailed experimental investigations, which will be presented in Part II of the paper. This leads to the conclusion that CDA airfoils that have been primarily developed for aeroengine applications are not the optimum solution, if directly transferred to heavy-duty gas turbines. A significant improvement in compressor efficiency is possible, if the new profiles are used instead of conventional airfoils.

Advanced High Turning Compressor Airfoils for Low Reynolds Number Condition: Part 1 — Design and Optimization

High performance compressor airfoils at a low Reynolds number condition at ~Re51.3310 5 ! have been developed using evolutionary algorithms in order to improve the performance of the outlet guide vane (OGV), used in a single low pressure turbine (LPT) of a small turbofan engine for business jet aircrafts. Two different numerical optimization methods, the evolution strategy (ES) and the multi-objective genetic algorithm (MOGA), were adopted for the design process to minimize the total pressure loss and the deviation angle at the design point at low Reynolds number condition. Especially, with respect to the MOGA, robustness against changes of the incidence angle is considered. The optimization process includes the representation of the blade geometry, the generation of a numerical grid and a blade-to-blade analysis using a quasi-three-dimensional Navier-Stokes solver with a k-v turbulence model including a newly implemented transition model to evaluate the performance. Overall aerodynamic performance and boundary layer properties for the two optimized blades are discussed numerically. The superior performance of the two optimized airfoils is demonstrated by a comparison with conventional controlled diffusion airfoils (CDA). The advantage in performance has been confirmed by detailed experimental investigations, which are presented in Part II of this paper. @DOI: 10.1115/1.1737780#

Effects of Reynolds Number and Free-Stream Turbulence on Boundary Layer Transition in a Compressor Cascade

An experimental and analytical study has been performed on the effect of Reynolds number and free-stream turbulence on boundary layer transition location on the suction surface of a controlled diffusion airfoil (CDA). The experiments were conducted in a rectilinear cascade facility at Reynolds numbers between 0.7 and 3.0 310 6 and turbulence intensities from about 0.7 to 4 percent. An oil streak technique and liquid crystal coatings were used to visualize the boundary layer state. For small turbulence levels and all Reynolds numbers tested, the accelerated front portion of the blade is laminar and transition occurs within a laminar separation bubble shortly after the maximum velocity near 35‐40 percent of chord. For high turbulence levels (Tu .3 percent) and high Reynolds numbers, the transition region moves upstream into the accelerated front portion of the CDA blade. For those conditions, the sensitivity to surface roughness increases considerably; at Tu54 percent, bypass transition is observed near 7 ‐10 percent of chord. Experimental results are compared to theoretical predictions using the transition model, which is implemented in the MISES code of Youngren and Drela. Overall, the results indicate that early bypass transition at high turbulence levels must alter the profile velocity distribution for compressor blades that are designed and optimized for high Reynolds numbers. @DOI: 10.1115/1.1413471#

Advanced High Turning Compressor Airfoils for Low Reynolds Number Condition—Part I: Design and Optimization

High performance compressor airfoils at a low Reynolds number condition at (Re = 1.3 × 10 5 ) have been developed using evolutionary algorithms in order to improve the performance of the outlet guide vane (OGV), used in a single low pressure turbine (LPT) of a small turbofan engine for business jet aircrafts. Two different numerical optimization methods, the evolution strategy (ES) and the multi-objective genetic algorithm (MOGA), were adopted for the design process to minimize the total pressure loss and the deviation angle at the design point at low Reynolds number condition. Especially, with respect to the MOGA, robustness against changes of the incidence angle is considered. The optimization process includes the representation of the blade geometry, the generation of a numerical grid and a blade-to-blade analysis using a quasi-three-dimensional Navier-Stokes solver with a k-ω turbulence model including a newly implemented transition model to evaluate the performance. Overall aerodynamic performance and boundary layer properties for the two optimized blades are discussed numerically. The superior performance of the two optimized airfoils is demonstrated by a comparison with conventional controlled diffusion airfoils (CDA). The advantage in performance has been confirmed by detailed experimental investigations, which are presented in Part II of this paper.

3D Transonic Flow in a Compressor Cascade With Shock-Induced Corner Stall

An experimental and numerical study of the transonic flow through a linear compressor cascade with endwalls was conducted. The cascade with a low aspect ratio of 1.34 was tested at an inlet Mach number of 1.09 and a Reynolds number of 1.9×106. Detailed flow visualizations on the surfaces and 5-hole probe measurements inside the blading and in the wake region showed clearly a 3-dimensional boundary layer separation on the blade surface and the sidewall, and a severe corner stall induced by a strong 3D shock system at blade passage entrance.The experimental data has been used to validate and improve the 3D Navier-Stokes code TRACE. Results showed an excellent resolution of the complex flow field. Surface pressure distributions on the entire blade surface and the endwalls, flow angle and total pressure contours within the blade passage and the wake are compared with the experimental results. An analysis of the secondary flow of this highly staggered cascade did not show the classical corner vortex. Instead, a severe flow deviation and partly reverse flow near the walls is seen. The flow solver helped to identify a weak ring vortex that originates from the passage sidewall. Surface oil flow pictures on the blade contour and the sidewall are in qualitatively good agreement to numerical surface streaklines. A considerable improvement of the numerical results could be achieved by a gradual grid refinement especially in the corner region and by successive code development.Copyright

1999 Turbomachinery Committee Best Paper Award: Development of Advanced Compressor Airfoils for Heavy-Duty Gas Turbines— Part II: Experimental and Theoretical Analysis

In Part I of this paper a family of numerically optimized subsonic compressor airfoils for heavy-duty gas turbines, covering a wide range of flow properties, is presented. The objective of the optimization was to create profiles with a wide low loss incidence range. Therefore, design point and off-design performance had to be considered in an objective function. The special flow conditions in large-scale gas turbines have been taken into account by performing the numerical optimization procedure at high Reynolds numbers and high turbulence levels. The objective of Part II is to examine some of the characteristics describing the new airfoils, as well as to prove the reliability of the design process and the flow solver applied. Therefore, some characteristic members of the new airfoil series have been extensively investigated in the cascade wind tunnel of DLR cologne. Experimental and numerical results show profile Mach number distributions, total pressure losses, flow turning, and static pressure rise for the entire incidence range. The design goal with low losses and especially a wide operating range could be confirmed, as well as a mild stall behavior. Boundary layer development, particularly near stall condition, is discussed using surface flow visualization and the results of boundary layer calculations. An additional experimental study, using liquid crystal coating, provides necessary information on suction surface boundary-layer transition at high Reynolds numbers. Finally, results of Navier-Stokes simulations are presented that enlighten the total pressure loss development and flow turning behavior, especially at high incidence in relation to the results of the design tool.

Advanced high-turning compressor airfoils for low Reynolds number condition: Part II: Experimental and numerical analysis

Part I of this paper describes the design and optimization of two high turning subsonic compressor cascades operating as an outlet guide vane (OGV) behind a single stage low pressure turbine at low Reynolds number condition (Re = 1.3 × 10 5 ). In the numerical optimization algorithm, the design point and off-design performance has been considered in an objective function to achieve a wide low loss incidence range. The objective of the present paper is to examine some of the characteristics describing the new airfoils as well as to prove the reliability of the design process and the applied flow salver Some aerodynamic characteristics for the two new airfoils and a conventional controlled diffusion airfoil (CDA), have been extensively investigated in the cascade wind tunnel of DLR Cologne. For an inlet Mach number of 0.6 the effect of Reynolds number and incidence angle on each airfoil performance is discussed, based on experimental and numerical results. For an interpretation of the airfoil boundary layer behavior, results of some boundary layer calculations are compared to oil flow visualization pictures. The design goal of an increased low loss incidence range at low Reynolds number condition could be confirmed without having a negative effect on the high Reynolds number region.

Advanced Nonaxisymmetric Endwall Contouring for Axial Compressors by Generating an Aerodynamic Separator—Part II: Experimental and Numerical Cascade Investigation

Modern methods for axial compressor design are capable of shaping the blade surfaces in a three-dimensional way. Linking these methods with automated optimization techniques provides a major benefit to the design process. The application of nonaxisymmetric contoured endwalls is considered to be very successful in turbine rotors and vanes. Concerning axial compressors, nonaxisymmetric endwalls are still a field of research. This two-part paper presents the recent development of a novel endwall design. A vortex created by a nonaxisymmetric endwall groove acts as an aerodynamic separator, preventing the passage vortex from interacting with the suction side boundary layer. This major impact on the secondary flow results in a significant loss reduction by means of load redistribution, reduction in recirculation areas, and suppressed corner separation. Part I of this paper deals with the endwall design and its compressor application. The resulting flow phenomena and physics are described and analyzed in detail. The second paper presents the detailed experimental and numerical investigation of the developed endwall groove. The measurements carried out at the transonic cascade wind tunnel of DLR in Cologne, demonstrated a considerable influence on the cascade performance. A loss reduction and redistribution of the cascade loading were achieved at the aerodynamic design point, as well as near the stall condition of the cascade. This behavior is well predicted by the numerical simulation. The combined analysis of experimental and numerical flow patterns allows a detailed interpretation and description of the resulting flow phenomena. In this context, high fidelity 3D-Reynolds-averaged Navier―Stokes flow simulations are required to analyze the complex blade and endwall boundary layer interaction.

3-D Transonic Flow in a Compressor Cascade With Shock-Induced Corner Stall

An experimental and numerical study of the transonic flow through a linear compressor cascade with endwalls was conducted. The cascade with a low aspect ratio of 1.34 was tested at an inlet Mach number of 1.09 and a Reynolds number of 1.9 310 6 . Detailed flow visualizations on the surfaces and five-hole probe measurements inside the blading and in the wake region showed clearly a three-dimensional boundary layer separation on the blade surface and the sidewall, and a severe corner stall induced by a strong 3-D shock system at blade passage entrance. The experimental data have been used to validate and improve the 3-D Navier-Stokes code TRACE. Results showed an excellent resolution of the complex flow field. Surface pressure distributions on the entire blade surface and the endwalls, flow angle and total pressure contours within the blade passage and the wake are compared with the experimental results. An analysis of the secondary flow of this highly staggered cascade did not show the classical corner vortex. Instead, a severe flow deviation and partly reverse flow near the walls is seen. The flow solver helped to identify a weak ring vortex that originates from the passage sidewall. Surface oil flow pictures on the blade contour and the sidewall are in qualitatively good agreement to numerical surface streaklines. A considerable improvement of the numerical results could be achieved by a gradual grid refinement, especially in the corner region and by successive code development.@DOI: 10.1115/1.1460913#

Aerodynamic Characteristics of Supercritical Outlet Guide Vanes at Low Reynolds Number Conditions

As a part of an innovative aerodynamic design concept for a single stage low pressure turbine, a high turning outlet guide vane is required to remove the swirl from the hot gas. The airfoil of the vane is a highly loaded compressor airfoil that has to operate at very low Reynolds numbers (Re∼ 120,000). Recently published numerical design studies and experimental analysis on alternatively designed airfoils showed that blade profiles with an extreme front loaded pressure distribution are advantageous for low Reynolds number conditions. The advantage even holds true for an increased inlet Mach number at which the peak Mach number on the airfoils reaches and exceeds the critical conditions (M ss >1.0). This paper discusses the effect of the inlet Mach number and Reynolds number on the cascade performance for both a controlled diffusion airfoil (CDA) (called baseline) and a numerically optimized front loaded airfoil. The results show that it is advantageous to design the profile with a fairly steep pressure gradient immediately at the front part in order to promote early transition or to prevent too large laminar-even shock induced-separations with the risk of a bubble burst. Profile Mach number distributions and wake traverse data are presented for design and off-design conditions. The discussion of Mach number distributions and boundary layer behavior is supported by numerical results obtained from the blade-to-blade flow solver MISES.