Skip to main content
Article thumbnail
Location of Repository

The effect of upstream duct boundary layer growth and compressor blade lean angle variation on an axial compressor performance

By I. Templalexis, Vassilios Pachidis, Pericles Pilidis and P. Kotsiopoulos


The compressor of a gas turbine engine is extremely vulnerable on upstream duct- induced flow non-uniformities whether the duct is an engine intake or an interconnecting duct. This is justified by its position being literally an extension of the duct flow path, coupled to the fact that it operates under adverse pressure gradients. In particular, this study focuses on performance deviations between installed and uninstalled compressors. Test results acquired from a test bed installation will differ from those recorded when the compressor operates as an integral part of an engine. The upstream duct, whether an engine intake or an inter-stage duct, will affect the flow-field pattern ingested into the compressor. The case study presented here aims mostly at qualifying the effect of boundary layer growth along the upstream duct wall on compressor performance. Additionally, the compressor performance response on blade lean angle variation is also addressed, with the aim of acquiring an understanding as to how compressor blade lean angle changes interact with intake-induced flow non-uniformities. Such studies are usually conducted as part of the preliminary design phase. Consequently, experimental performance investigation is excluded at this stage of development, and therefore, computer-aided simulation techniques are used if not the only option for compressor performance prediction. Given the fact that many such design parameters need to be assessed under the time pressure exerted by the tight compressor development programme, the compressor flow simulation technique needs to provide reliable results while consuming the least possible computational time. Such a low computational time compressor flow simulation method, among others, is the two-dimensional streamline curvature (SLC) method, being also applied within the frame of reference of the current study. The paper is introduced by a brief discussion on SLC method. Then, a reference is made to the radial equilibrium equation, which is the mathematical basis of SOCRATES, a turbomachinery flow simulation tool that was used in this study. Subsequently, the influence of the upstream duct on the compressor inlet radial flow distribution is being addressed, with the aim of adjusting the compressor blade inlet lean angle, in order to minimize compressor performance deterioration. The paper concludes with a discussion of the results

Topics: lean angle, intake, streamline curvature, radial equilibrium equation, compressor performance
Publisher: Professional Engineering Publishing
Year: 2010
DOI identifier: 10.1243/09544100JAERO637
OAI identifier:
Provided by: Cranfield CERES

Suggested articles


  1. (1958). A Practical Method Of Predicting Transonic-Compressor Performance”, doi
  2. (1957). A preliminary analysis of the magnitude of shock losses in transonic compressors”,
  3. (2006). Abudus Samad, Kwang-Yong Kim “Optimal Design of Swept, Leaned and Skewed Blades in a Transonic Axial Compressor”, GT2006-90384, ASME Turbo Expo, Power For Land, Sea and Air, doi
  4. An Iterative Method for Blade Profile Loss Model Adaptation Using Streamline Curvature”, doi
  5. (1949). Application Of Radial Equilibrium Condition To Axial-flow Compressor And Turbine Design”,
  6. (2007). ASME Cycle Innovations Committee Best Paper Award for
  7. (1968). Axial Flow Compressor Computer Program for Calculating Off-Design Performance (Program IV)”, General Motors, Allison Division, Indianapolis, Prepared for NASA,
  8. (2003). Axial-Flow Compressors: A Strategy For Aerodynamic Design And Analysis”, doi
  9. (1965). Chapter VI - Experimental Flow In
  10. Compressor With Distorted Inlet Flow.”, doi
  11. (2008). Development of a 2D compressor streamline curvature code”, doi
  12. (1991). Inlet Distortion Effects In Aircraft Propulsion System Integration”.
  13. (1977). Performance And Stability Of A J85-
  14. Performance Of Two-Stage Fan Having Low-AspectRatio, First Stage Rotor Blading”,
  15. (2007). Prediction of Engine Performance Under Compressor Inlet Flow Distortion Using Streamline Curvature”, doi
  16. (1970). Secondary flow losses in axial compressors.”
  17. (1961). Shock Losses In Transonic Rotor Rows”, doi
  18. (1950). The Low Speed Performance Of Related Aerofoils In Cascade”,
  19. (2008). Turbo Expo, Power For Land, Sea and Air,
  20. (2002). Xu “The Effects of Lean and Sweep on Transonic Fan Performance”, GT2002-30327, ASME Turbo Expo, Power For Land, Sea and Air, doi

To submit an update or takedown request for this paper, please submit an Update/Correction/Removal Request.