In-Flight Visualization of Supersonic Flow Transition Using Infrared Imaging

Dam, C. P. van; Shiu, Henry; Banks, David C.; Tracy, Richard; Chase, Jared M.

doi:10.2514/2.3046

Cited by 4 publications

(5 citation statements)

References 5 publications

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“…A train of swept shocks appear to be affecting the transition mechanism by inducing a spanwise pressure gradient similar to that found on swept wings. The origin of this train of shocks in unclear; it was not observed on the previous NLF flight tests [15], though it is apparent that the smaller test article described in van Dam et al had its leading edge behind the swept shock train. Tests conducted on the flat plate did not show the shocks, but the test article had little insulation and the thermal effect of the shocks on the surface could have been diffused on the highly conductive aluminum substrate [19] …”

Section: Figure 6 Evolution Of Transition Location With Altitude and mentioning

confidence: 93%

“…In 1999 NASA and Aerion's founders worked together in testing the technology development of supersonic NLF using the F-15B supersonic testbed, [15] where an NLF test article demonstrated laminar flow under a complex flow field up to 7.8 million Reynolds number. Other tests using different means have been used in demonstrating natural laminar in supersonic flight to extents up to 30 million Reynolds number and documented in [16,17,18] In the last few years natural laminar flow robustness has been an important part of Aerion's NLF research as part of its path to implement natural laminar flow on a supersonic civil aircraft.…”

Section: Introductionmentioning

confidence: 99%

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Supersonic NLF Robustness Flight Testing: Transition Due to Discrete Roughness Elements

Garzon¹,

Matisheck²,

Banks

et al. 2014

32nd AIAA Applied Aerodynamics Conference

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This paper describes flight test results of a large-scale supersonic natural laminar flow airfoil. These tests were conducted primarily in the spring of 2013. The test article was 79 inches in chord and 40 inches in span and was mounted on the centerline store location of an F-15B testbed aircraft. Supersonic tests were conducted from 31,000 ft to 49,500 ft altitudes and at Mach numbers up to 2.0. Boundary layer transition was measured using an aircraft mounted infrared camera system. In addition to baseline transition characteristics a main objective of this test series was to determine the sensitivity of the boundary layer to discrete disturbances. The results show that even with some unwanted flow interactions, apparently from the test fixture, the airfoil performed well. In addition, the results indicate the boundary layer sensitivity to disturbances, and thus future manufacturing tolerances, were similar to that of subsonic natural laminar flow airfoils. NomenclatureD = Diameter or characteristic length of the roughness element k = Roughness element height H = Flight altitude in ft M = Mach number Re = Freestream Reynolds number per unit length Rek = Reynolds number based on local flow conditions at top of roughness and roughness height Re = Reynolds number based on local boundary layer parameters and momentum thickness T = Temperature  = Boundary layer momentum thickness Subscripts e = Physical property or condition measured or evaluated at the boundary layer edge k = Physical property or condition measured or evaluated at the roughness height tr = transition w = Physical property or condition measured or evaluated at the wall

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Section: Figure 6 Evolution Of Transition Location With Altitude and mentioning

confidence: 93%

Section: Introductionmentioning

confidence: 99%

Supersonic NLF Robustness Flight Testing: Transition Due to Discrete Roughness Elements

Garzon¹,

Matisheck²,

Banks

et al. 2014

32nd AIAA Applied Aerodynamics Conference

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“…The test is described in more detail by van Dam et al 11 The test article featured a biconvex airfoil with a leading edge sweep of 15 degrees, and a trailing edge angle of 30 degrees. It was fabricated of aluminum and covered with rigid closed cell foam and an outer layer of epoxy.…”

Section: A Flight Testsmentioning

confidence: 99%

Supersonic Testing of Natural Laminar Flow on Sharp Leading Edge Airfoils. Recent Experiments by Aerion Corporation

Garzon¹,

Matisheck²

2012

42nd AIAA Fluid Dynamics Conference and Exhibit

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Though inviscid drag can be the biggest component of drag on supersonic aircraft, viscous drag reduction can account for a major improvement in total drag. A significant reduction in viscous drag can be achieved by extensive laminar flow on the wings, which can be obtained by carefully designing the airfoil's pressure gradients. This tailoring of the design is more difficult if supercritical airfoils in conjunction with high leading edge sweep are adopted. In such case the extent of laminar flow is reduced due to either adverse pressure gradient early in the boundary layer development and/or growth of cross flow instabilities. A way of avoiding these issues is to unsweep the wing and adopt biconvex-type sharp leading edge airfoils. This concept has been tested and adopted by Aerion Corporation and test results showing extensive runs of natural laminar flow (NLF), both in ground facilities and in flight, are summarized.

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Section: Introductionmentioning

confidence: 99%

“…The SSNLF test article included an insulative coating material to minimize heat conduction to the aluminum substructure. Results from the SSNLF project demonstrated the feasibility of maintaining large runs of laminar flow on low-swept wings at chord Reynolds numbers up to 10 million [11,12]. An IR image obtained from one of the SSNLF flights The Supersonic Boundary Layer Transition (SBLT) project followed the SSNLF testing with the objective of expanding the results from SSNLF to more complex surfaces with larger chord lengths (higher chord Reynolds numbers).…”

Section: Introductionmentioning

confidence: 99%

Flight tests of a supersonic natural laminar flow airfoil

Frederick¹,

Banks²,

Garzon³

et al. 2015

Meas. Sci. Technol.

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A flight test campaign of a supersonic natural laminar flow airfoil has been recently completed. The test surface was an 80-inch (203 cm) chord and 40-inch (102 cm) span article mounted on the centerline store location of an F-15B airplane. The wing was designed with a leading edge sweep of effectively 0 deg to minimize boundary layer crossflow. The test article surface was coated with an insulating material to avoid significant heat transfer to and from the test article structure to maintain a quasi-adiabatic wall. An aircraft-mounted infrared camera system was used to determine boundary layer transition and the extent of laminar flow. The tests were flown up to Mach 2.0 and chord Reynolds numbers in excess of 30 million. The objectives of the tests were to determine the extent of laminar flow at high Reynolds numbers and to determine the sensitivity of the flow to disturbances. Both discrete (trip dots) and 2-D disturbances (forward-facing steps) were tested. A series of oblique shocks, of yet unknown origin, appeared on the surface, which generated sufficient crossflow to affect transition. Despite the unwanted crossflow, the airfoil performed well. The results indicate the sensitivity of the flow to the disturbances, which can translate into manufacturing tolerances, were similar to that of subsonic natural laminar flow wings.

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In-Flight Visualization of Supersonic Flow Transition Using Infrared Imaging

Cited by 4 publications

References 5 publications

Supersonic NLF Robustness Flight Testing: Transition Due to Discrete Roughness Elements

Supersonic NLF Robustness Flight Testing: Transition Due to Discrete Roughness Elements

Supersonic Testing of Natural Laminar Flow on Sharp Leading Edge Airfoils. Recent Experiments by Aerion Corporation

Flight tests of a supersonic natural laminar flow airfoil

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