The effect of flow confinement on laminar shock-wave/boundary-layer interactions

Lusher, David J.; Sandham, Neil D.

doi:10.1017/jfm.2020.354

Cited by 23 publications

(24 citation statements)

References 44 publications

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“…Since skin-friction drag is the major source of drag (contributing about half of the total aircraft drag [ 4 , 5 ]), substantial friction drag reduction can be achieved by maintaining the flow laminar over large portions of the aircraft surfaces [ 5 , 6 ]. However, laminar boundary layers are also more prone to separation than their turbulent counterparts [ 7 , 8 ]. Flow separation may have a negative impact on the aerodynamic performance of the aircraft surfaces, especially when induced by a Shock-Wave/Boundary-Layer Interaction (SWBLI) [ 7 , 8 , 9 , 10 , 11 ].…”

Section: Introductionmentioning

confidence: 99%

“…However, laminar boundary layers are also more prone to separation than their turbulent counterparts [ 7 , 8 ]. Flow separation may have a negative impact on the aerodynamic performance of the aircraft surfaces, especially when induced by a Shock-Wave/Boundary-Layer Interaction (SWBLI) [ 7 , 8 , 9 , 10 , 11 ]. This complex phenomenon [ 11 , 12 , 13 ] is likely to occur at transonic flow conditions on aircraft wings designed with laminar flow technology, where the laminar boundary layer over the suction side of the wing reaches supersonic speeds; the supersonic flow region is typically terminated by a (normal) shock, which interacts with the boundary layer and induces flow separation [ 1 , 14 , 15 ].…”

Section: Introductionmentioning

confidence: 99%

“…In this case, the flow generally undergoes transition to turbulence, and the strongly increased wall-normal transport of momentum (and energy) eventually leads to the reattachment of the turbulent flow to the surface. The resulting region of reverse flow enclosed within the boundary layer is commonly referred to as a Laminar Separation Bubble (LSB) [ 7 , 8 , 10 , 14 , 16 , 17 ]. Laminar SWBLI can also occur in other aerodynamic applications, including internal flows in the gas turbine engines of transonic commercial aircraft [ 7 , 18 , 19 ] and in the engines of supersonic and hypersonic vehicles [ 8 , 10 , 16 ].…”

Section: Introductionmentioning

confidence: 99%

“…The resulting region of reverse flow enclosed within the boundary layer is commonly referred to as a Laminar Separation Bubble (LSB) [ 7 , 8 , 10 , 14 , 16 , 17 ]. Laminar SWBLI can also occur in other aerodynamic applications, including internal flows in the gas turbine engines of transonic commercial aircraft [ 7 , 18 , 19 ] and in the engines of supersonic and hypersonic vehicles [ 8 , 10 , 16 ]. Major concerns arise in the presence of laminar SWBLI because it can cause detrimental effects for the aircraft performance, such as drag increase, flow unsteadiness, local heat peaks, high aerodynamic loads and increased structural fatigue of the aircraft components [ 7 , 8 , 9 , 10 , 16 , 19 ].…”

Section: Introductionmentioning

confidence: 99%

“…Laminar SWBLI can also occur in other aerodynamic applications, including internal flows in the gas turbine engines of transonic commercial aircraft [ 7 , 18 , 19 ] and in the engines of supersonic and hypersonic vehicles [ 8 , 10 , 16 ]. Major concerns arise in the presence of laminar SWBLI because it can cause detrimental effects for the aircraft performance, such as drag increase, flow unsteadiness, local heat peaks, high aerodynamic loads and increased structural fatigue of the aircraft components [ 7 , 8 , 9 , 10 , 16 , 19 ]. From these considerations, it appears clear that accurate information on the occurrence of a laminar separation bubble induced by SWBLI, including the locations of flow separation and reattachment over the whole examined surface, is crucial in a variety of aerodynamic applications.…”

Section: Introductionmentioning

confidence: 99%

See 4 more Smart Citations

Skin-Friction-Based Identification of the Critical Lines in a Transonic, High Reynolds Number Flow via Temperature-Sensitive Paint

Costantini

Henne

Klein

et al. 2021

Sensors

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In this contribution, three methodologies based on temperature-sensitive paint (TSP) data were further developed and applied for the optical determination of the critical locations of flow separation and reattachment in compressible, high Reynolds number flows. The methodologies rely on skin-friction extraction approaches developed for low-speed flows, which were adapted in this work to study flow separation and reattachment in the presence of shockwave/boundary-layer interaction. In a first approach, skin-friction topological maps were obtained from time-averaged surface temperature distributions, thus enabling the identification of the critical lines as converging and diverging skin-friction lines. In the other two approaches, the critical lines were identified from the maps of the propagation celerity of temperature perturbations, which were determined from time-resolved TSP data. The experiments were conducted at a freestream Mach number of 0.72 and a chord Reynolds number of 9.7 million in the Transonic Wind Tunnel Göttingen on a VA-2 supercritical airfoil model, which was equipped with two exchangeable TSP modules specifically designed for transonic, high Reynolds number tests. The separation and reattachment lines identified via the three different TSP-based approaches were shown to be in mutual agreement, and were also found to be in agreement with reference experimental and numerical data.

show abstract

Section: Introductionmentioning

confidence: 99%

Section: Introductionmentioning

confidence: 99%

Section: Introductionmentioning

confidence: 99%

Section: Introductionmentioning

confidence: 99%

Section: Introductionmentioning

confidence: 99%

See 3 more Smart Citations

Skin-Friction-Based Identification of the Critical Lines in a Transonic, High Reynolds Number Flow via Temperature-Sensitive Paint

Costantini

Henne

Klein

et al. 2021

Sensors

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Shockwave/Boundary-Layer Interactions in Transitional Rectangular Duct Flows

Lusher

Sandham

2020

ERCOFTAC Series

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Shock-wave/boundary-layer interactions (SBLI) are an important feature of high-speed gas dynamics. In many numerical studies of SBLI span-periodicity is assumed to reduce computational complexity. However, span-periodicity is not a valid assumption for aeronautical applications such as supersonic engine intakes where lateral confinement leads to highly three-dimensional behaviour. In this work transitional oblique SBLI are simulated for a rectangular duct with a θ sg = 5 • shock generator ramp at Mach 2. The baseline configuration is a duct with an aspect ratio of 0.5. Time-dependent disturbances are added to the base laminar flow via wall localised blowing/suction strips to obtain intermittent transition upstream of the SBLI. Two forcing configurations are evaluated to assess the response of the SBLI to different tripping locations. The transition is observed to develop first in the low-momentum corners of the duct and spread out in a wedge shape. The central separation bubble is seen to react dynamically to oncoming turbulent spots, shifting laterally across the span. While instantaneous corner separations do occur, the time-averaged corner flow remains attached. Comparisons to a one-to-one aspect ratio duct show that the SBLI is heavily dependent on aspect ratio; the wider duct exhibited significantly larger regions of flow-reversal due to a strengthened interaction.

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Entrainment Mechanism Analysis of Oblique Shock-Wave/Boundary-Layer Interactions

Meng,

Han,

Yang

2024

IUTAM Bookseries

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The interaction between shock waves and boundary layers impacts hypersonic vehicles’ aerodynamic structure, internal turbulent mixing, and combustion processes. Direct numerical simulations (DNSs) of the hypersonic boundary layer have been performed to study the turbulent/non-turbulent interface (TNTI) entrainment mechanism and the impact of shock waves. Two cases are analyzed with and without the shock-wave and boundary layer interaction. A novel approach is proposed to identify TNTI by integrating the vorticity threshold and fuzzy clustering method. To account for the influence of increasing Reynolds number in the flow direction on the vorticity threshold analysis, vorticity is normalized using the local boundary layer thickness and friction velocity. Two entrainment mechanisms are quantitatively described and compared using TNTI local entrainment velocity and the mass conservation equation. The results demonstrate that the entrainment process within the hypersonic boundary layer TNTI is predominantly governed by large-scale engulfment, and under the influence of shock waves, the dominance of large-scale ingestion is notably enhanced.

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The effect of flow confinement on laminar shock-wave/boundary-layer interactions

Cited by 23 publications

References 44 publications

Skin-Friction-Based Identification of the Critical Lines in a Transonic, High Reynolds Number Flow via Temperature-Sensitive Paint

Skin-Friction-Based Identification of the Critical Lines in a Transonic, High Reynolds Number Flow via Temperature-Sensitive Paint

Shockwave/Boundary-Layer Interactions in Transitional Rectangular Duct Flows

Entrainment Mechanism Analysis of Oblique Shock-Wave/Boundary-Layer Interactions

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