Effect of Reynolds number on a normal shock wave-transitional boundary-layer interaction over a curved surface

Coschignano, A; Atkins, Nicholas R.; Babinsky, Holger; Serna, Juana

doi:10.1007/s00348-019-2824-0

Cited by 10 publications

(6 citation statements)

References 17 publications

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“…The experiment runs at a Reynolds number relative to the nacelle lip thickness of 1.2 × 10 6 while the full-scale nacelle operates The FLIR SC7300LW set to an integration time of 320 µs is used for IR visualisation. Previous work with IR in the blow-down tunnel by Coschignano et al (2019) meant that a suitable IR window and camera mount is available, allowing for camera set up in less than an hour. The FLIR A615 was also tested and found to be sensitive enough for visualisation in the experiment.…”

Section: Case Studymentioning

confidence: 99%

Infrared Thermography Techniques for Boundary Layer State Visualisation

Davis,

Atkins

2023

Preprint

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The rapid decarbonisation of the power generation and aviation sectors will require a move away from incremental development, exposing designers and researchers to the risk of unexpected results from uncertainty in boundary layer state. This problem already exists for parts developed with fully turbulent assumptions, but in novel design spaces the risk increases for both real components, where previous knowledge of similar designs may be inapplicable, and particularly in experimental testing of scaled models, where reducing Reynolds number can result in a drastic change in flow topology that skews the conclusions of a test. Computational methods struggle to reliably predict boundary layer state so experimental techniques for diagnosing boundary layer state are needed. Infrared thermography (IR) is a non-invasive technique that offers simple, fast visualisation of boundary layer state with no additional instrumentation. IR is relatively uncommon in the literature and there is minimal information available on the best practices for its use. This paper aims to encourage the adoption of IR as a diagnostic tool by demonstrating routes for optimisation and pointing out pitfalls to avoid. A low-order model is developed and used to predict how the signal-to-noise ratio (SNR) of an IR visualisation changes depending on the thermal design of the test piece. It is shown that in low-speed flows, with active heating from the surface, the SNR is maximised through a suitable choice of surface insulation while in high-speed flows, where passive temperature differences are used, there is a crossover between heat transfer and recovery temperature effects that results in an SNR of zero, an effect that can arise in both steady-state and transient experiments. Experimental validation of the 1D model in both flow regimes is shown alongside two case studies on the use of IR in sub-scale testing where uncertainty in boundary layer state results in critical differences from the full-scale flow.

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Section: Case Studymentioning

confidence: 99%

Infrared Thermography Techniques for Boundary Layer State Visualisation

Davis,

Atkins

2023

Preprint

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“…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 ]. 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.…”

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%

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.

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“…Further complicated phenomena are also observed by the presence of the transonic flow such as the interaction between -shock and the boundary layer flow [1][2][3]. These phenomena relate to flow separation, thus causing a negative effect on lift production of the wing [4][5]. The ONERA M6 wing, which was experimentally reported by Schmitt and Charpin [6], is a typical example of a transonic wing that can perform the mentioned phenomena, as presented by Dwight [7].…”

Section: Introductionmentioning

confidence: 99%

Parametric Study on Wing-Lambda-Shock Formation

Chainok

Rungroch

Chairach

et al. 2021

Volume 1: Aerospace Engineering Division Joint Track; Computational Fluid Dynamics

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It is well-known that a wing is one of the most important parts of an aircraft as it is used to generate lift force. According to a wing moving at sufficiently high subsonic speeds, the flow speed on the wing’s upper surface can be supersonic due to acceleration through the curvature-created suction, thereby forming a shock wave in a lambda shape. Additionally, the lambda shock can interact with the boundary layer flow. These phenomena relate to disturbances in the flow field, including flow separation, thus causing undesirable effects on lift production. Hence, a better understanding of the phenomenon of wing-lambda-shock formation and its nature is essential. This study presents a numerical investigation of the lambda-shock formation on an ONERA M6 wing, which is known as a swept, semi-span wing with no twist, under parametric effects of angle-of-attack, and free-stream Mach number, which is increased up to the supersonic regime. The pressure coefficients obtained by simulations are validated by open data. Then, numerical results in terms of the local pressure coefficient, local Mach number, averaged lift and drag coefficients, and λ-shape characteristics based on Mach number and pressure coefficients are discussed under an investigated range of the parameters. Results show that the angle-of-attack and free-stream Mach number can affect the lambda shock formation on the wing upper surface physically. Specifically, an iso-sonic surface with lambda shock waves is disturbed when the angle-of-attack and free-stream Mach number vary in an investigated range. This also affects lift and drag coefficients of the wing.

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Effect of Reynolds number on a normal shock wave-transitional boundary-layer interaction over a curved surface

Cited by 10 publications

References 17 publications

Infrared Thermography Techniques for Boundary Layer State Visualisation

Infrared Thermography Techniques for Boundary Layer State Visualisation

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

Parametric Study on Wing-Lambda-Shock Formation

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