The Calculation of the Three-Dimensional Flow Through a Transonic Fan Including the Effects of Blade Surface Boundary Layers, Part-Span Shroud, Engine Splitter and Adjacent Blade Rows PUBLIC ACCESS

[+] Author Affiliations
R. D. Cedar

GE Aircraft Engines, Cincinnati, OH

D. G. Holmes

GE Research and Development Center, Schenectady, NY

Paper No. 89-GT-325, pp. V001T01A114; 9 pages
  • ASME 1989 International Gas Turbine and Aeroengine Congress and Exposition
  • Volume 1: Turbomachinery
  • Toronto, Ontario, Canada, June 4–8, 1989
  • Conference Sponsors: International Gas Turbine Institute
  • ISBN: 978-0-7918-7913-9
  • Copyright © 1989 by ASME


In the past, predictions of the three-dimensional flow through turbomachinery blades have often assumed that:-

a) The blade passage forms a single channel.

b) The effects of adjacent blade rows and engine structures are negligible.

c) The flow is inviscid.

Unfortunately, for a transonic fan in a high bypass ratio aero-engine, these assumptions are poor because:-

a) The blade passage is split into a number of channels by the part-span shroud and the engine splitter.

b) The fan aerodynamics may be influenced by the presence of the engine splitter and downstream blade rows.

c) The blockage due to the blade surface boundary layers can have a significant effect on the shock structure.

This paper describes the extensions that have been made to the three-dimensional Euler solver developed by Holmes and Tong (1985) in order to remove the assumptions listed above. The problem of modeling the part-span shroud and the engine splitter has been generalized to that of solving a flow containing any number of internal solid bodies. This has been achieved in a novel manner that allows the finite volume scheme to be fully vectorized and the multigrid acceleration technique to be implemented efficiently. The effect of adjacent blade rows has been modeled using passage averaged mass, momentum and energy source terms to represent the neighboring blades in a similar manner to that described by Adamczyk, Mulac and Celestina (1986). The blade surface boundary layers have been modeled using a series of two-dimensional boundary layer calculations coupled to the Euler solver using a transpiration model to account for the boundary layer displacement thickness. To demonstrate the capabilities of this program, results from the analysis of a General Electric advanced fan blade design, with a part-span shroud, engine splitter and down-stream stator and rotor blade are presented. This is followed by an investigation and discussion of the individual effect of not modeling the viscous blockage, the part-span shroud engine, splitter and down-stream blade rows. It is concluded that if an accurate prediction of the fan aerodynamics is to be obtained all these features must be modeled.

Copyright © 1989 by ASME
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