This paper presents results of numerical investigations carried out to explore the benefit of endwall boundary layer removal from critical regions of highly-loaded axial compressor blade rows. At the loading level of modern aero-engine compressors the performance is primarily determined by three-dimensional flow phenomena occuring in the endwall regions. 3DNS simulations were conducted on both a rotor and a stator test case in order to evaluate basic effects and the practical value of bleeding air from specific locations at the casing endwall. The results of the numerical survey demonstrated substantial benefits of relatively small bleed rates to the local flow field and to the performance of the two blade rows. On the rotor, boundary layer fluid was removed from the main flow path through an axisymmetric slot in the casing over the rotor tip. This proved to give some control over the tip leakage vortex flow and the associated loss generation. On the stator, boundary layer fluid was taken from the flow path through a single bleed hole within the passage. Two alternative off-take configurations were evaluated, revealing a large impact of the bleed hole shape and location on the cross-passage flow and the suction side corner separation. On both blade rows investigated, rotor and stator, boundary layer removal resulted in a reduction of local reverse flow, blockage and losses in the respective near-casing region. This paper gives insight into changes occuring in the 3D passage flow field near the casing and summarises the effects on the radial matching and pitchwise-averaged performance parameters, namely loss and deviation of the rotor and stator when suction is active. Primary focus is put on the aerodynamics in the blade rows in the main flow path; details of the internal flow structure within the bleed off-take cavities/ports are not discussed here.
Stable Combustion of a High-Velocity Gas in a Heated Boundary Layer
