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Old 10-09-2020, 12:08 PM   #1 (permalink)
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You-Tube : NACA airfoil, smoke-flow imaging

Last Wednesday, one of the members shared a link to a video, from NACA ( now NASA ), depicting an NACA 0012 in a wind tunnel, undergoing smoke flow visualization at varying degrees of angle of attack.
I couldn't find it here this morning, and I won't be back until maybe Monday next week, so I'm going to post some information in case the author needs it in the mean time.
--------------------------------------------------------------------------------------
Presented randomly:
* data from the NACA Variable Density Tunnel ( VDT )
* the 0012 designates an ' symmetrical' airfoil, of ' 12-percent' thickness ratio.
* rectangular airfoil
* aspect ratio = 6.0
* velocity = 80-feet / second
* air density ( density altitude )
* normal surface roughness
* ambient airstream turbulence
* Reynolds number approx. 3,500,000
* NACA also presents values @ Rn= 6,000,000, and Rn = 9,000,000, with and without angled flaps
* section reaction forces are Reynolds number-dependent
* aerodynamic center at 30% chord, from leading edge
* laminar boundary layer exists up to 30% chord ( 1st minimum pressure )
* from 30%-to-100% chord, boundary layer is turbulent
* specific weight of air = 32.2 rho
* turbulence approx. 0.025 ( 2.5 % )
* airfoil reaction forces are proportional to air density
* airfoil reaction forces are proportional to 'true' speed
* zero lift, CL min = zero @ 0-degrees AOA ( it's symmetrical )
* Cdo min = 0.0083 @ 0-degrees AOA ( it's symmetrical )
* @ 0-degrees AOA, the 0012 section has no net negative pressure over the last 20% of chord.
* CL max = 1.53 @ 22-degrees AOA ( W. A. Mair's magic 22-degrees )
* beyond 22-degrees, adverse pressure gradient overwhelms the boundary layers ability to remain attached, with ensuing separation, and turbulence
* Cd & CL are constant for a given AOA, regardless of altitude
* Low and slow drag = High and fast drag
* Center of pressure / chord = Center of Pressure Coefficient
* Chord is the geometric chord, from forward stagnation point, through rear stagnation point
* Span = wing span
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* with a traditional Clark Y section, stall occurs at 25-degrees AOA.( If one has enough altitude, simply take your hands and feet off all controls, and the airplane will recover straight and level flight )
* adding a front slot to the Clark Y increases stall angle to 35-degrees
* with blown or suctioned slots, a Clark Y can withstand even greater AOA without stall
*Mean chord = surface area ( S ) divided by ( b ) span
* Aspect ratio ( AR ) = Span / mean chord
* Burble-point / Stall = flow separation over the upper surface
* Moment coefficients act around the Aerodynamic Center ( a.c. )
* any wing section may have up to four ( 4) different performance curves, based upon :
1) 'normal' section
2) same section with flap
3) same section with slot
4) same section with both flap and slot


* The airplane must fly faster to support itself at altitude due to the lower air density
* Drag is less at altitude

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Last edited by aerohead; 10-09-2020 at 12:09 PM.. Reason: typo
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Old 10-10-2020, 10:27 AM   #2 (permalink)
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One of my favourite airfoils only seconded with the NACA 64Xxx series but without the annoying abrupt stall

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