Two-dimensional chord-wise load distributions at transonic speeds Page: 4 of 43
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NACA RM L51107
t - thickness at any chordwise station.
max maximum value
APPARATUS AND TESTS
The tests were conducted in the Langley 4- by 19-inch semiopen
tunnel (fig. 1), which is an induction tunnel housed within an enclosure
to minimize condensation problems. (Enclosure discussed in reference 5.)
The 4-inch dimension of the tunnel is formed by two parallel steel plates,
whereas the 19-inch-dimension is the dimension of the open jet, the top
and bottom chambers of which are connected by a duct. Velocity distri-
butions in the test region obtained from static-pressure measurements
at the tunnel walls are shown in figure 2. Figure 2(a) shows that the
velocities along the longitudinal axis at the model location are within
1 percent for stream Mach numbers around 0.8. For stream Mach numbers
near 1.0, the velocities are within 2 percent. The velocity gradients
along the normal-to-chord axis (fig. 2(b)) are smaller.
Each airfoil completely spanned the test section along the 4-inch
dimension of the tunnel and was supported by large circular end plates
which fitted into the tunnel walls and rotated with the model to retain
continuity of the surfaces of the tunnel walls. The juncture between
airfoil and tunnel wall was sealed.
The investigation' included schlieren photographs of the flow and
static-pressure-distribution measurements on the airfoils. Approximately
40 static-pressure orifices were installed in two rows 0.25 inch on
either side of the center line of the models. Test runs were made at
constant angles of attack through a Mach number range. The stream Mach
number was controlled by regulating the mass flow through the tunnel
by using a sonic-throat device, designated "choker" in figure 1, located
downstream of the test section.
Variations in airfoil thickness were investigated by using the
NACA 64A-series profile, varying in thickness from 4 to 12 percent of
chord (fig. 3(a)). Effects of thickness distribution were investigated
by tests on the NACA 6A- and 16-series profiles having a constant thick-
ness of 9 percent (fig. 3(b)). The extent of the investigation on camber
was limited to tests on 6-percent-thick airfoils of the NACA 64A series
varying in camber from zero to a design lift coefficient of 0.5
(fig. 3(c)). The test Reynolds number of these 4-inch-chord airfoils
at a Mach number of 1.0 was 1.6 X 106.
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Lindsey, Walter F. & Dick, Richard S. Two-dimensional chord-wise load distributions at transonic speeds, report, February 18, 1952; (digital.library.unt.edu/ark:/67531/metadc58975/m1/4/: accessed November 21, 2018), University of North Texas Libraries, Digital Library, digital.library.unt.edu; crediting UNT Libraries Government Documents Department.