Summary of Airfoil Data Page: 18 of 266
[265] p. : ill.View a full description of this report.
Extracted Text
The following text was automatically extracted from the image on this page using optical character recognition software:
REPORT NO. 824 -NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
the MIach numbers at which large changes in airfoil char-
acteristics occur, especially when sharp pressure peaks exist
at the leading edge. A discussion of the characteristics of
airfoil sections at supercritical M ach numbers is beyond the
scope of this report.
For convenience, curves of predicted critical Mach num-
ber plotted against the low-speed section lift coefficient have
been included in the supplementary figures for a number of
airfoils. High-speed lift coefficients may be obtained by
multiplying the low-speed lift coefficient by the factor
1
1-M The critical Iach numbers have been predicted
from theoretical pressure distributions. For airfoils of the
NACA four- and five-digit series and for the NACA 7-series
airfoils, the theoretical pressure distributions were obtained
by Theodorsen's method. For the other airfoils the theo-
retical pressure distributions were obtained by the approxi-
mate method describedd in the preceding section.
The data in the supplementary figures show that, for any
one type of airfoil, the maximum critical Iach number
decreases rapidly as the thickness is increased. The effect
of camber is to lower the maximum critical Mach number
and to shift the range of high critical Mach numbers in the
same manner as for the low drag range. For common types
of camber the minimum reduction in critical speed for a
given design lift coefficient is obtained with a uniform load
type of mean line. A comparison of the data presented in
the supplementary figures shows that NACA 6-series sec-
tions have considerably higher maximum critical Malch
numbers than NACA 24-, 44-, and 230-series airfoils of
corresponding thickness ratios.
MOMENT COEFFICIENTS
Methods of calculation.-Theoretical moment coefficients
may be approximated directly from the values presented in
the supplementary figures for the various mean lines. These
values were obtained from thin-airfoil theory and may be
scaled up or down linearly with the design lift coefficient or
with the mean-line ordinates. These theoretical values are
sufficiently accurate for preliminary considerations, but ex-
perimental values should be used for stability and control
calculations.
Numerical examples.-The following numelrical examples
illustrate the methods of calculating the moment coefficients:
Example 1: Find the theoretical moment coefficient about
thel quarter-chor(l point for thle NACA 652 215, a= 0.5
airfoil.
'hj (lesigllatiol of the' airfoil shows that thle de(I(sign lift
(oeflicicnt of tids airfoil is 0.2. From the data on the
NACA ~t= ().5 typeL mean line inl ieI n tlhe, suiplem entairy
figures, tlie value of c,,,, is -0.139 for a (lesign lift coeffi(iceIt
of 1.0. The desiredd value of the moimieint (of(dil tnt is
C,,, (-0.1 39) (0.2)
-- -0.028Exanl)l(' 2: Find the theoretical moment coeffici ent about
the (uarter-ch'lord point for the N ACA 4415 airfoil.
From the descriptionn of the NACA four-digit series
airfoils, the required data is found to be presellted( for theNACA 64 mean line in the supplementary figures. The
moment coefficient for this mean line is -0.1W7. The
required value is then
mc-(-0.157)
S-0.105
ANGLE OF ZERO LIFT
Methods of calculation.-Values of the ideal or design
angle of attack as corresponding to the design lift coefficient
c, are included among the data for the various mean lines
presented in the supplementary figures. The approximate
values of the angle of zero lift may be obtained from the
data by using the theoretical value of the lift-curve slope
for thin airfoils, 2w per radian. The value of ato in degrees
is then57.3
oa= C-a- 2 C-c(16)
The tabulated values of a, may be scaled linearly with
the design lift coefficient or with the mean-line ordinates.
Although these theoretical angles of zero lift may be useful
in preliminary (design, they should not be used without
experimental verification for such purposes as establishing
the washout of a wing.
Numerical examples.-The method of computing a0 is
illustrated in the following examples:
Example 1: Find the theoretical angle of zero lift of the
NACA 652 -515, a= 0.5 airfoil.
This airfoil number indicates a design lift coefficient of
0.5. Data for the NACA a=0.5 mean line indi('ate that
a= 3.040 when c-,=1.0. The desired value of a, is then
a,i=(3.04) (0.5)
-1.52
Substituting in equation (16) gives
a 0=J52-(57.3) (0.5)
-3.0
Example 2: Find the theoretical angle of zero lift for the
NACA 2415 airfoil.
The description of the NACA four-digit-series airfoils
shows that the required values of ai and cz, may be obtained
by multiplying the corresponding values for the NACA 64
mean line (see supplementary figures) by a factor 2/6; then
ac= (0.74)
=0.25
= 0.253
an(l from equation (16)
a, 0.2 (57.3) (0.253)
2.0014
Upcoming Pages
Here’s what’s next.
Search Inside
This report can be searched. Note: Results may vary based on the legibility of text within the document.
Tools / Downloads
Get a copy of this page or view the extracted text.
Citing and Sharing
Basic information for referencing this web page. We also provide extended guidance on usage rights, references, copying or embedding.
Reference the current page of this Report.
Abbott, Ira H.; von Doenhoff, Albert E. & Stivers, Louis S., Jr. Summary of Airfoil Data, report, 1945; (https://digital.library.unt.edu/ark:/67531/metadc65534/m1/18/: accessed May 11, 2024), University of North Texas Libraries, UNT Digital Library, https://digital.library.unt.edu; crediting UNT Libraries Government Documents Department.