Low-speed Linear Compressor Cascade Research Articles

Experimental Investigations on the Effects of Surface Roughness for Compressor Cascades With Different Roughness Magnitude and Location

Abstract Blade surface roughness could significantly affect the aerodynamic performance of compressors. To explore the influence of roughness magnitude and location on blade performance, experiments were conducted in a low-speed linear compressor cascade with controlled diffusion airfoils (CDA). A part-span roughness method was employed in the experiment to maintain the axial velocity–density ratio (AVDR) during the change of blade roughness magnitudes and locations. Five blade surface local roughness schemes, including the leading-edge, the fore- and aft-part of the suction surface, and the pressure surface, which were determined based on geometry sensitivity analysis, were investigated with the variation of the surface roughness magnitude between Ra = 3.1 μm to 18.8 μm. Cascade inlet and outlet flowfields and the blade surface static pressure distributions were measured, which could help to distinguish the change of blade performance characteristics and even blade surface boundary layer development state. A critical roughness effect was found, and significant blade loss increment and available incidence range reduction appear at super-critical roughness states. At the measured maxi-roughness condition, 28.4% loss increase and 41.2% incidence range reduction were reached.

Impact of Surface Roughness on Compressor Cascade Performance

This paper presents an experimental investigation of roughness effects on aerodynamic performance in a low-speed linear compressor cascade. Equivalent sandgrain roughnesses of 12 μm, 180 μm, 300 μm, 425 μm, and 850 μm have been tested. In nondimensional terms, these roughnesses represent compressor blade roughnesses found in actual gas turbines. Downstream pressure, velocity, and angle have been measured with a five-hole probe at 0.3 chord downstream of the blade trailing edge. For the tested roughnesses of 180 μm, 300 μm, 425 μm, and 850 μm, the axial velocity ratio across the blade row decreases by 0.1%, 2.1%, 2.5%, and 5.4%, respectively. For the same cases, the exit flow angle deviation increases by 24%, 38%, 51%, and 70%, respectively. Finally, the mass-averaged total pressure loss increases by 12%, 44%, 132%, and 217%, respectively. Also, the loss increases more rapidly in the transitionally rough region. Thus, among the three parameters, the loss responds most sensitively to changes in compressor blade roughness.

Wake of a Compressor Cascade with Tip Gap, Part 3: Two Point Statistics

Velocity spectra and space-time correlations have been measured downstream of a low-speed linear compressor cascade with tip gap. The objective of this work is an improved understanding of the coherent turbulence structures present in fan-tip wakes and of the turbulent source terms responsible for broadband stator noise

Numerical Simulation of Three-Dimensional Viscous Flow in a Linear Compressor Cascade With Tip Clearance

A Navier–Stokes solver is applied to investigate the three-dimensional viscous flow in a low-speed linear compressor cascade with tip clearance at design and off-design conditions with two different meshes. The algebraic turbulence model of Baldwin–Lomax is used for closure. Relative motion between the blades and wall is simulated for one flow coefficient. Comparisons with experimental data, including flow structure, static and total pressures, velocity profiles, secondary flows and vorticity, are presented for the stationary wall case. It is shown that the code predicts well the flow structure observed in experiments and shows the details of the tip leakage flow and the leading edge horseshoe vortex.