Fixed sum airloads on an airfoil

Hello,

Reading books on aerodynamic/design/conception/construction of aircraft gives you
(by date of pub.) by authors
Munk
Abbott (and related NACA 814 reports)
Peery
Roskam
Bruhn
Raymer (on design)
Snorri Gudmundsson
and a lot of others authors (naming here Americans ones). I haven't read
all, but some (or part of it).

In France we have also old literature and old engineering regulation (that's
so different from now for example giving moment coefficient in wind tunnel
at leading edge when NACA was giving them at 25% (aerodynamic center) chords).

But old literature about old regulation is giving sometimes good and relevant
numbers for homebuilt airplanes.
In immediate post WWII age, in France was a regulation Air2004B that some french
authors illustrated in a rather easy way.

A pitfall in conception is the travel of center of pressure
at different Cm/Cl that made the NACA 23012 a wide accepted airfoil that has
Cm in 0.01 range meaning that center of pressure is almost near the aerodynamic
center(no aft 30%). Harry Riblett gives you a bunch of information on this airfoil
that is prone to icing (on ATR family of aircraft) and exhibits an abrupt stall
that is the heads of the (no center of pressure traveling) tails.

Correction : H.Riblett writes about the NACA 43018 airfoil for ATR family
of aircraft.
Airfoil 230XX and 430XX are depicted in NACA report N°610 related to Forward camber
airfoils. For Airfoil 43018 Cm is also in the -0.01 range (cf p.19 of report)


Air2004B regulation gave 4 different fixed sum airloads on an airfoil depending
on center of pressure location i.e Cm/Cl

The "weird" thing is CoG of surfaces gives you (algebraically) the position
of the center of pressure, and sum of airloads on airfoil.


I tried to find the same calculus in different books and didn't found anything
that resemble it. If you have seen something close to this in a textbook or have
any commentary on it, you're welcome to share it.


This post "Fixed Sum Airloads on an airfoil" is for educational purposes only and may contain errors. Any attempt
to use the results for actual design purposes are done at the user's own risk.

My previous post was just plain wrong; formulas were for a
center of pressure under 1/3 chord, such as ironically
NACA 23000 series of airfoil.


For center of pressure over 1/3 chord the fixed sum
airloads on airfoil is as illustrated below.


The center of pressure at Cl 0.2 (0.717 chord) is
far to aft and over the wing chord that's make
the calculus at this point erroneous.


So if you have an idea for that (Cl 0.2 limit
and acceptable center of pressure/wing chord
ratio) you're welcome to share it.

This post "Fixed Sum Airloads on an airfoil" is for educational purposes only and may contain errors. Any attempt
to use the results for actual design purposes are done at the user's own risk.