Application of Advanced Computational Codes in the Design of an Experiment for a Supersonic Throughflow Fan Rotor

Abstract

Increased emphasis on sustained supersonic or hypersonic cruise has revived interest in the supersonic throughflow fan as a possible component in advanced propulsion systems. Use of a fan that can operate with a supersonic inlet axial Mach number is attractive from the standpoint of reducing the inlet losses incurred in diffusing the flow from a supersonic flight Mach number to a subsonic one at the fan face. The data base for components of this type is practically nonexistent; therefore, in order to furnish the required information for assessment of this type fan, a program has been initiated at the NASA Lewis Research Center to design, build, and test a fan rotor that operates with supersonic axial velocities from inlet to exit. This paper describes the design of the experiment using advanced computational codes to calculate the unique components required. The fan rotor has constant hub and tip radii and was designed for a pressure ratio of 2.7 with a tip speed of 457 m/s. The rotor was designed using existing turbomachinery design and analysis codes modified to handle fully supersonic axial flow through the rotor. A two-dimensional axisymmetric throughflow design code plus a blade element code were used to generate fan rotor velocity diagrams and blade shapes. A quasi-three-dimensional, thin shear layer Navier–Stokes code was used to assess the performance of the fan rotor blade shapes. The final design was stacked and checked for three-dimensional effects using a three-dimensional Euler code interactively coupled with a two-dimensional boundary layer code. A translating nozzle was designed to produce a uniform flow parallel to the fan up to the design axial Mach number of 2.0. The nozzle was designed with the three-dimensional Euler/interactive boundary layer code. The nozzle design in the expansion region was analyzed with a three-dimensional parabolized viscous code, which corroborated the results from the Euler code. A translating supersonic diffuser was designed using these same codes.

Supersonic Axial Flow Oscillating Cascade Shock Motion and Unsteady Aerodynamic Moment

“…W = ΠΑΜΙ (4) Thus, the unsteady aerodynamic work per cycle is directly proportional to the out-of-phase component of the unsteady moment. This is the work done by the air on the airfoil.…”

A modal aeroelastic finite element analysis method for advanced turbomachinery stages

“…Assuming a linear variation of the loading between two strips directly results in (15). To further increase the method's flexibility, (16) is transformed in order to function with basic section motions. For each strip h, modal displacements ( O p h ) are known and thus can be approximated easily by the combination of a few elementary section motions, which can be calculated using a mean-squares polynomial approximation method.…”

The aerodynamics of a baseline supersonic throughflow fan rotor

“…Based on a personal survey of the literature, it appears that there are only two docu mented cases prior to 1989 involving successful experiments where axial-supersonic flow into a rotating blade row was established, and in both of these cases only shock-in-rotor operation was examined, as was probably intended. Indeed, prior to recent efforts at the NASA Lewis Research Center [27][28][29][30][31] there appears to have been no reported experimen tal work aimed at directly investigating a supersonic throughflow rotor ^ or stage, although detailed design studies [32,33] were done in the mid 1960's which included experimental research on a supersonic throughflow rotor cascade [34]. Linear cascade tests with axialsupersonic flows into or through blade passages are not particularly scarce.…”

“…The basic approach for achieving the program goals involves conducting, in parallel, experimental and analytical research so as to take advantage of the strengths inherent in each. A somewhat "conservative" baseline proof-of-concept fan design has served as the starting point for the program, where heavy reliance on advanced computational fluid dynamics (CFD) codes was necessary since no relevant experimental data base existed for such designs [27,28]. Details of the baseline fan design and the basic design philosophy used are reviewed in the next chapter.…”

“…The baseline SSTF was designed in-house at the NASA Lewis Research Center in 1986. The design is reviewed in this chapter, which contains information previously reported in two NASA Technical Memorandums [27,28]. Only a summary is presented here since complete details can be found in the references.…”