# Review of Passive Flow Control Methods for Compressor Linear Cascades
## 2 Chronological Evolution of Linear Cascade
### Passive Control Methods
#### 2.1.3 Passive Flow Control via Surface Modifications  
The role of boundary layer behaviour in compressor performance has been a focus of research, with studies on controlled diffusion airfoils (CDA) showing that higher turbulence levels move the transition point upstream. In 2014, suction at 70% of the axial chord improved boundary layer conditions and reduced passage total pressure losses [132,133]. Variations in inlet flow angle, Mach number, and Reynolds number (450,000 to 900,000) significantly affected total pressure loss. Shock-induced transition occurred over separated flow at lower Reynolds numbers, and without separation at higher Reynolds numbers [134]. Further research into controlling corner flow in a highly loaded axial compressor cascade has shown that endwall suction, particularly using a pitchwise slot, provides the best performance across various incidence angles. Compound suction configurations, combining a spanwise blade slot with a pitchwise endwall slot, proved even more effective, achieving a total pressure loss reduction of up to 39.4% at 7\({}^{\circ}\) incidence. For accurately capturing the complex 3D separations in the transonic cascade, the Reynolds Stress Model (RSM) was identified as the most suitable turbulence model [135]. The impact of blade end slots and full-span slots on corner separation was also evaluated. Slotted profiles suppressed separation, particularly at high incidence angles, with blade end slots outperforming full-span slots by reducing mixing losses at mid-span [136].  
In 2021, rotor-stator interactions were found to significantly influence secondary flow patterns and boundary layer transitions. Shock Control Bump (SCB) technology was applied to transonic compressor blades, effectively reducing aerodynamic losses and improving efficiency at peak speeds [137; 138]. Figure 9 presents the geometry of this type of shock control bump, where the initial position of the SCB structure is crucial to suppress loss of stagnation pressure in blade passages. The SCB structure is composed of ramp, crest plateau, and tail sections, connected smoothly as a curve with third-order continuity using fourth-order spline fitting [139].  
Further investigations on shock wave-boundary layer interactions using perforated plates and porous wall designs revealed that these methods reduced shock intensity and improved efficiency while managing flow separation [139; 140]. Figure 10 illustrates a perforated airfoil (main and secondary blades) used to suppress the shock waves of a centrifugal compressor diffuser.  
Additionally, porous bleed holes in transonic compressor cascades resulted in up to a 75% reduction in total pressure losses by delaying flow separation through better boundary layer control [141]. A study on bleed pressure in transonic compressors demonstrated that lower bleed pressure ratios notably reduced pressure losses. The changes in shock wave structure led to the creation of favourable pressure gradients, which in turn enhanced boundary layer stability [142]. Furthermore, an investigation into shock-wave/boundary-layer interactions in transonic compressor cascades showed that increasing Mach numbers had a substantial effect on shock patterns. The presence of separation bubbles and shock oscillations resulted in significant fluctuations in the pressure field [143].