Nonadiabatic Surface Effects on Transition Measurements Using Temperature-Sensitive Paints

Costantini, Marco; Fey, Uwe; Henne, Ulrich; Klein, Christian

doi:10.2514/1.j053155

Cited by 34 publications

(34 citation statements)

References 27 publications

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“…Subsequently, this lower temperature remains approximately constant during the actual run for some 10ths of 1 s, whereas the model temperature drifts just slightly from its prerun values. When the prerun model temperature and the charge temperature of the gas are approximately the same, this leads to a temperature difference between flow and the model surface during a run [24,26,30]. In this case, called in the following the "standard DNW-KRG thermal condition at the model surface," the ratio between the model surface temperature and the adiabatic-wall temperature T w ∕T aw is significantly larger than one.…”

Section: Cryogenic Ludwieg-tube Göttingenmentioning

confidence: 99%

“…At subsonic Mach numbers, the axial pressure gradient was moderately favorable in the wind-tunnel tests and essentially zero in the flight tests, with this difference being due to the influence of the fuselage in the flight tests. In [24], an NLF airfoil was experimentally investigated at both low and high subsonic Mach numbers for varying angles of attack. The results showed that the impact of a nonadiabatic wall on boundary layer transition depended on the boundary layer stability situation.…”

Section: Introductionmentioning

confidence: 99%

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Pressure Gradient and Nonadiabatic Surface Effects on Boundary Layer Transition

et al. 2016

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The influence of the streamwise pressure gradient and a nonadiabatic surface on boundary layer transition was experimentally investigated at the Cryogenic Ludwieg-Tube Göttingen, Germany. Boundary layer transition was detected nonintrusively by means of the temperature-sensitive paint technique. The wind-tunnel model was designed to achieve a quasi-uniform streamwise pressure gradient over a large portion of the model chord length. This allowed the effects on boundary layer transition of the streamwise pressure gradient and wall temperature ratio to be decoupled. The model was tested at high Reynolds numbers and at a high subsonic Mach number. Favorable, almostzero, and adverse streamwise pressure gradients were considered; and various temperature differences between the flow and the model surface were implemented. Stronger flow acceleration and lower wall temperature ratios led to an increase of the transition Reynolds number. Larger increases in the transition Reynolds number were obtained at more pronounced flow acceleration for the same reduction in the wall temperature ratio. The measured pressure distributions served as input for boundary layer stability computations, performed according to compressible, linear, local stability theory under the (quasi-) parallel-flow assumption. Amplification factors of the Tollmien-Schlichting waves were shown to be reduced by stronger favorable pressure gradients and lower wall temperature ratios, which was in agreement with the observed variation in the transition Reynolds number.

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Section: Cryogenic Ludwieg-tube Göttingenmentioning

confidence: 99%

Section: Introductionmentioning

confidence: 99%

Pressure Gradient and Nonadiabatic Surface Effects on Boundary Layer Transition

et al. 2016

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“…Heat flux between the surface and flow is different for laminar and turbulent flows. As mentioned, in low Re flows, LSB forms due to the laminar flow separation and turbulent flow reattachment therefore detecting these areas utilizing IT allows the opportunity of detecting LSB characteristics on the surface of the airfoil [8]. A useful study related to separation detection using IT was reported [9] where the flow about a cylinder was studied and also investigated the separated flow area in deep stall condition over an airfoil using thermographic imaging, and temperature fluctuation to detect the separated flow.…”

Section: Introductionmentioning

confidence: 99%

Evaluating airfoil behaviour such as laminar separation bubbles with visualization and IR thermography methods

Ghorbanishohrat

Johnson

2018

J. Phys.: Conf. Ser.

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Abstract. A study of airfoil behaviour is undertaken using two independent non-intrusive measurement techniques, surface oil flow visualization (SOFV) and infrared thermography (IT) to quantitatively determine flow parameters such as separation and reattachment points on the airfoil. Several airfoil heating methods are described and several airfoil materials are evaluated. Results are presented here for constant Re = 4.7 × 10 4 , with the angle of attack varied from 0 • to 13 • . Innovative post processing methods are used for the IR temperature data to reveal important details of the flow. All methods provide comparable results and allow further understanding of these characteristics and the extension to large scale, dynamic and unsteady flow situations.

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“…Recent studies by Constantini et al [29,30] have shown that care must be taken when interpreting transition results obtained using TSP and applying a temperature step. Their work has shown that when the surface temperature (T w ) is greater than the adiabatic–wall temperature (T aw ), then the transition location can vary depending on the T w /T aw ratio.…”

Section: Resultsmentioning

confidence: 99%

“…Their work has shown that when the surface temperature (T w ) is greater than the adiabatic–wall temperature (T aw ), then the transition location can vary depending on the T w /T aw ratio. To see if the application of different voltages to the CNT layer can cause a change in the transition position, an analysis of the transition location was carried out for each case using a similar procedure outlined by Constantini et al [29] and the results are shown in Figure 9. For this work, the transition position when 50 V was applied was difficult to visualize accurately.…”

Section: Resultsmentioning

confidence: 99%

Boundary-Layer Detection at Cryogenic Conditions Using Temperature Sensitive Paint Coupled with a Carbon Nanotube Heating Layer

Goodman

Lipford

Watkins

2016

Sensors

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Detection of flow transition on aircraft surfaces and models can be vital to the development of future vehicles and computational methods for evaluating vehicle concepts. In testing at ambient conditions, IR thermography is ideal for this measurement. However, for higher Reynolds number testing, cryogenic facilities are often used, in which IR thermography is difficult to employ. In these facilities, temperature sensitive paint is an alternative with a temperature step introduced to enhance the natural temperature change from transition. Traditional methods for inducing the temperature step by changing the liquid nitrogen injection rate often change the tunnel conditions. Recent work has shown that adding a layer consisting of carbon nanotubes to the surface can be used to impart a temperature step on the model surface with little change in the operating conditions. Unfortunately, this system physically degraded at 130 K and lost heating capability. This paper describes a modification of this technique enabling operation down to at least 77 K, well below the temperature reached in cryogenic facilities. This is possible because the CNT layer is in a polyurethane binder. This was tested on a Natural Laminar Flow model in a cryogenic facility and transition detection was successfully visualized at conditions from 200 K to 110 K. Results were also compared with the traditional temperature step method.

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Nonadiabatic Surface Effects on Transition Measurements Using Temperature-Sensitive Paints

Cited by 34 publications

References 27 publications

Pressure Gradient and Nonadiabatic Surface Effects on Boundary Layer Transition

Pressure Gradient and Nonadiabatic Surface Effects on Boundary Layer Transition

Evaluating airfoil behaviour such as laminar separation bubbles with visualization and IR thermography methods

Boundary-Layer Detection at Cryogenic Conditions Using Temperature Sensitive Paint Coupled with a Carbon Nanotube Heating Layer

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