Application of sweep to improve the efficiency of a transonic fan. I - Design

Abstract

The concept of aerodynamic sweep has been applied to the design of a highly swept, low-aspect ratio, transonic fan rotor. This design, to be described here, represents the second phase of the Navy Advanced Fan Component Technology program aimed at quantifying the combined performance benefits of low-aspect ratio and of leading- edge aerodynamic sweep on high-pressure ratio transonic fan stages. The first phase included a design and test of a baseline low-aspect ratio unswept fan stage having a design total pressure ratio of 2.2, a tip speed of 1558 ft/s, and relative tip Mach number of 1.6. This baseline fan, designed and tested in 1986-87, achieved its design goals. The design of the swept rotor with the same application requirements was also completed in 1986. The swept design was predicted to be shock free and to improve fan rotor adiabatic efficiency 1.5%. The swept fan performance testing was completed in 1988, and intrablade velocities were measured in 1989. ERODYNAMIC sweep has been applied to aircraft wing design since the 1940s to reduce drag at transonic and supersonic flight velocities. It has also been apparent since then that aerodynamic sweep principles might be applicable to turbomachiner y blading having supersonic relative Mach number to achieve significant improvements in efficiency. Early attempts were made from the 1940s to 1970s to apply sweep to turbomachinery blading, but were not successful due to the difficulties involved in translating this technology into the confined rotating environment of the gas turbine engine. In this environment, the fan inlet relative Mach number in- creases from hub to tip due to rotation, both ends of the airfoil are confined, neighboring airfoils are present, and the objective is to achieve a static pressure rise. This technology development culminated in the NASA-sponsored QF-12 fan stage design and test.1'3 The rotor in this stage had a design supersonic inlet relative tip Mach number of 1.588 and com- pound swept blading that reduced the effective relative Mach number to subsonic values, ranging from 0.83 at the hub to 0.91 at the mean and tip. The primary goal of the design was to reduce acoustic noise relative to a conventional design through the reduction of shock wave generated multiple pure tones. Unfortunately, the fan rotor failed to achieve the aero- dynamic design goals of flow, pressure ratio, and efficiency. Some acoustic tone noise improvements were measured, but again, these changes were below the design expectations. These deficiencies were apparently due to the unexpected radial distribution of the inlet flow induced by the swept rotor. It was also concluded that the potential improvements could be reached with future designs. The recent development of three-dimensi onal computa- tional fluid dynamic analysis for transonic turbomachinery blade rows and stages has made the achievement of these envisioned performance improvements feasible in the 1990s. In a recent NASA-sponsored program, these new computa- tional methods were used to explore the application of sweep to moderate pressure ratio, high-aspect ratio fan stages,4 and in a subsequent Navy-sponsored Navy Advanced Fan Com- ponent Technology (NAFCOT) program this new method was used to design a high-pressure ratio, low-aspect ratio fan stage with both an unswept 5 and a swept rotor. This stage includes inlet guide vanes (IGV) and exit stators. The swept design will be described here. The swept rotor performance test program and intrablade laser velocity results are found in Ref. 6, and further details are found in the contract report.7