Unsteadiness of shock-boundary layer interactions in a Mach 2.0 supersonic turbine cascade

Lui, Hugo F. S.; Ricciardi, Túlio Rodarte; Wolf, William; Braun, James; Rahbari, Iman; Paniagua, Guillermo

doi:10.1103/physrevfluids.7.094602

Cited by 9 publications

(8 citation statements)

References 66 publications

Supporting

Mentioning

Contrasting

Order By: Relevance

“…Wall-resolved LES are performed to study the interaction between shock waves and turbulent boundary layers in a linear cascade of supersonic turbines. The vane geometry is the same used by Liu et al [49] and Lui et al [3], as well as the fluid properties and inflow conditions, but here we vary the Mach number by specifying a uniform inlet flow at three different Mach numbers, M ∞ = 1.85, 2.00, and 2.15. The Reynolds number based on the inlet velocity and axial airfoil chord c x is fixed at Re = 395000.…”

Section: Flow Configuration and Numerical Methodologymentioning

confidence: 99%

“…To introduce minimal numerical dissipation in the vicinity of the shock waves without damping the small scales of turbulence, the localized artificial diffusivity (LAD) scheme LAD-D2-0 [52] is used to compute extra artificial bulk viscosity and thermal conductivity, which are then added to their physical counterparts. More details about the governing equations and numerical procedure for the LES are provided by Nagarajan et al [50], Bhaskaran and Lele [53], and Lui et al [3]. Supersonic inflow is used as the inlet boundary condition, while at the outflow, the Navier-Stokes characteristic boundary condition (NSCBC) is employed [54].…”

Section: Flow Configuration and Numerical Methodologymentioning

confidence: 99%

“…Although turbine blades are exposed to inflow turbulence, here we consider a uniform inlet flow but a body-forcing term is used to mimic a tripping device and induce transition to turbulence on the airfoil boundary layers [3,55]. This allows for a more direct comparison with other SBLI studies in the literature.…”

Section: Flow Configuration and Numerical Methodologymentioning

confidence: 99%

“…However, curved walls are encountered in some high-speed aerodynamic devices, such as turbomachinery blades, transonic airfoils, and nozzles. Some authors have investigated the unsteadiness of SBLIs on curved surfaces [3,[44][45][46], but most of these studies are related to transonic flows, which differ in multiple aspects from supersonic configurations given that, in the former, acoustic waves can travel upstream and affect the SBLIs [33]. Recently, Lui et al [3] performed a wall-resolved large eddy simulation (LES) of a supersonic turbine cascade at Mach 2.0.…”

mentioning

confidence: 99%

“…Some authors have investigated the unsteadiness of SBLIs on curved surfaces [3,[44][45][46], but most of these studies are related to transonic flows, which differ in multiple aspects from supersonic configurations given that, in the former, acoustic waves can travel upstream and affect the SBLIs [33]. Recently, Lui et al [3] performed a wall-resolved large eddy simulation (LES) of a supersonic turbine cascade at Mach 2.0. The curved walls of the airfoil led to distinctive shock structures compared to flat plates.…”

mentioning

confidence: 99%

See 4 more Smart Citations

Mach number effects on shock-boundary layer interactions over curved surfaces of supersonic turbine cascades

Lui,

Wolf,

Ricciardi

et al. 2024

Preprint

View full text Add to dashboard Cite

The effects of inlet Mach number on the unsteadiness of shock-boundary layer interactions (SBLIs) over curved surfaces are investigated for a supersonic turbine cascade using wall-resolved large eddy simulations. Three inlet Mach numbers, 1.85, 2.00, and 2.15 are considered at a chord-based Reynolds number 395000. The curved walls of the airfoils impact the SBLIs due to the state of the incoming boundary layers and local pressure gradients. On the suction side, due to the convex wall, the boundary layer entering the SBLI evolves under a favorable pressure gradient and bulk dilatation. On the other hand, the concave wall on the pressure side imposes an adverse pressure gradient and bulk compression. Variations in the inlet Mach number induce different shock impingement locations, enhancing these effects. A detailed characterization of the suction side boundary layers indicates that a higher Mach number leads to larger shape factors, favoring separation and larger bubbles, while the reverse holds for the pressure side. A time-frequency analysis reveals the presence of intermittent events in the separated flow occurring predominantly at low-frequencies on the suction side and at mid-frequencies on the pressure side. Increasing the inlet Mach number leads to an increase in the time scales of the intermittent events on the suction side, which are associated with instants when high-speed streaks penetrate the bubble, causing local flow reattachment and bubble contractions. Instantaneous flow visualizations show the presence of streamwise vortices developing on the turbulent boundary layers on both airfoil sides and along the bubbles. These vortices influence the formation of the large-scale longitudinal structures in the boundary layers, affecting the mass imbalance inside the separation bubbles.

show abstract

Section: Flow Configuration and Numerical Methodologymentioning

confidence: 99%