EcoModder Forum You-Tube : NACA airfoil, smoke-flow imaging

Register Now
 Remember

 10-09-2020, 12:08 PM #1 (permalink) Master EcoModder     Join Date: Jan 2008 Location: Sanger,Texas,U.S.A. Posts: 12,445 Thanks: 20,007 Thanked 6,216 Times in 3,837 Posts 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 ------------------------------------------------------------------------------------ * 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 __________________ Photobucket album: http://s1271.photobucket.com/albums/jj622/aerohead2/ Last edited by aerohead; 10-09-2020 at 12:09 PM.. Reason: typo
 The Following User Says Thank You to aerohead For This Useful Post: Piotrsko (10-10-2020)
 Today Popular topics Other popular topics in this forum...
 10-10-2020, 10:27 AM #2 (permalink) Somewhat crazed     Join Date: Sep 2013 Location: 1826 miles WSW of Normal Posts: 1,746 Thanks: 173 Thanked 483 Times in 412 Posts One of my favourite airfoils only seconded with the NACA 64Xxx series but without the annoying abrupt stall
 The Following User Says Thank You to Piotrsko For This Useful Post: aerohead (10-12-2020)