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 geometrical 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 which 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 which have been primarily developed for aero engine 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.

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 3.0×106 and turbulence intensities from about 0.7 to 4%. 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% of chord. For high turbulence levels (Tu > 3%) and high Reynolds numbers transition propagates upstream into the accelerated front portion of the CDA blade. For those conditions, the sensitivity to surface roughness increases considerably and at Tu = 4% bypass transition is observed near 7–10% 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.

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 10 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 windtunnel 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 which enlighten the total pressure loss development and flow turning behavior especially at high incidence in relation to the results of the design tool.

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 3.0×106 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 Tu=4 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.

No abstract