Stator well flows in axial compressors

Öztürk, Harun Kemal; Turner, A. B.; Childs, Peter; Bayley, F. J.

doi:10.1016/s0142-727x(00)00071-0

Cited by 5 publications

(2 citation statements)

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“…The tip vortex formation further forms as blockage in the flow field and this blockage can be defined in several ways. One of the well-known definition given by Kenneth L. Suder [5] is that it is the effective reduction in the flow area and is denoted by 'B' shown in the Equation (6). Equation (7) represents the mathematics representation of blockage 'B', Where A is the area of rotor blade.…”

Section: Understanding the Flow Domain Boundary Layer Of The Compressmentioning

confidence: 99%

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Flow blockage in a transonic axial flow compressor: simulation analysis under distorted conditions

Srinivas

Raghunandana

Shenoy

2018

IJET

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Today the aircraft industry is looking for faster and safer engines for both civil and military applications. The performance of all different types of air breathing engines depends on the amount of mass flow rate of air entering and hot gas ejecting out from the engine. Thrust is the key role for any engine performance. To achieve more thrust all the turbo machinery components like axial fan, axial flow compressor and axial flow turbine should function effectively. This paper is primarily dealing about one of the turbo machinery component, axial flow compressor performance where the study is more focused on flow blockage formation under distorted phenomena. The complete blade boundary layer formation and related flow numerical theory are discussed in detail, accordingly the boundary conditions were set to have better numerical simulations using ANSYS tool. To find the flow blockage formation suitable turbulence model was coded using the well know compressible equations. The flow blockage between the rotor and stator of the compressor stage was calculated and also validated with that of experimental data effectively. The flow simulation results also revealed that the performance parameters under the modern engine transonic speed from Mach 0.8 to 1.2 under the distorted conditions are better for aeromechanical features.

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Section: Understanding the Flow Domain Boundary Layer Of The Compressmentioning

confidence: 99%

“…H.K. Ozturk et al [6] discussed the design of the stator blade in a modern high-pressure turbofan engine. The purpose of this paper is to show turbulence effects, without flow blockage, especially seal leakage, can be reduced by incorporating fins on the rotating component rotor blade.…”

Section: Introductionmentioning

confidence: 99%