Effects of Reynolds number and leading-edge roughness on lift and drag characteristics of the NACA 65₃-418, a = 1.0 airf...

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Material Information

Title:
Effects of Reynolds number and leading-edge roughness on lift and drag characteristics of the NACA 65₃-418, a = 1.0 airfoil section
Alternate Title:
NACA wartime reports
Physical Description:
7, 8 p. : ill. ; 28 cm.
Language:
English
Creator:
Quinn, John H
Langley Aeronautical Laboratory
United States -- National Advisory Committee for Aeronautics
Publisher:
Langley Memorial Aeronautical Laboratory
Place of Publication:
Langley Field, VA
Publication Date:

Subjects

Subjects / Keywords:
Reynolds number   ( lcsh )
Aerodynamics -- Research   ( lcsh )
Genre:
federal government publication   ( marcgt )
bibliography   ( marcgt )
technical report   ( marcgt )
non-fiction   ( marcgt )

Notes

Summary:
Summary: Tests were made in the Langley two-dimensional low-turbulence tunnels of an NACA 65₃-418, a = 1.0 airfoil section with roughness in the form of carborundum grains applied to the leading edge. Roughness grains having average diameters of 0.0003 and 0.0007 airfoil chord were applied to the leading edge of the wing, and lift and drag measurements were made for a range of Reynolds numbers from 0.23 to 3.0 x 10⁶. From a comparison of data obtained in the present tests with data obtained in tests of the smooth wing, marked reductions in maximum lift coefficient were found to be caused by the roughness throughout the test range of Reynolds number. The drag coefficient at the design lift coefficient increased sharply and the lift-curve slope decreased rapidly at a critical Reynolds number that depended upon the size of the carborundum grains. This critical Reynolds number occurred at approximately 0.50 and 0.70 x 10⁶ for the 0.0003- and the 0.0007-chord-diameter roughness grains, respectively. With roughness, a decrease in maximum lift coefficient as great as 0.2, a decrease in lift-curve slope of 0.028, and an increase in drag coefficient at the design lift coefficient of 0.0007 were observed at a Reynolds number of 1.0 x 10⁶. For the smooth wing at the same Reynolds number, the maximum lift coefficient was 1.19, the lift-curve slope was 0.116, and the drag coefficient was 0.0077. At Reynolds numbers greater than 1.0 x 10⁶, the scale effect on the lift and drag characteristics of the section with both degrees of roughness was generally in the same direction as the effect on the lift and drag characteristics of the smooth airfoil.
Bibliography:
Includes bibliographic references (p. 7).
Statement of Responsibility:
by John H. Quinn, Jr.
General Note:
"Report no. L-82."
General Note:
"Originally issued November 1945 as Confidential Bulletin L5J04."
General Note:
"NACA WARTIME REPORTS are reprints of papers originally issued to provide rapid distribution of advance research results to an authorized group requiring them for the war effort. They were previously held under a security status but are now unclassified. Some of these reports were not technically edited. All have been reproduced without change in order to expedite general distribution."

Record Information

Source Institution:
University of Florida
Rights Management:
All applicable rights reserved by the source institution and holding location.
Resource Identifier:
aleph - 003638084
oclc - 71822444
sobekcm - AA00006273_00001
System ID:
AA00006273:00001

Full Text


CB No. L5JO0


NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS





WA11RTIME REPORT
ORIGINALLY ISSUED
November 1945 as
Confidential Bulletin L5JO4

EFFECTS OF REYNOLDS NUMBER AND LEADING-EDGE ROUGICiESS
ON LIFT AND DRAG CHARACTERISTICS OF THE
NACA 653-418, a = 1.0 AIRFOIL SECTION
By John H. Quinn, Jr.

Langley Memorial Aeronautical Laboratory
Langley Field, Va.







.... ACA$


WASHINGTON
NACA WARTIME REPORTS are reprints of papers originally issued to provide rapid distribution of
advance research results to an authorized group requiring them for the war effort. They were pre-
viously held under a security status but are now unclassified. Some of these reports were not tech-
nically edited. All have been reproduced without change in order to expedite general distribution.


L 82


DOCUMENTS DEPARTMENT




























































Ij; !1'!'








NACA CB No. L5.TOL

NATIONAL ADVISORY COMMITTEE FOR AERONAJUTICS


C'NFI DETIAL BULLETIN

EFFECTS OF REYNOLDS IUM3IBER AND LEADING-EDGE ROUGHNESS

Oil LIFT AND DRAG CHARACTERISTICS OF THE

I ACA 65z-11, a = 1.0 AIRFOIL SECTION

By John H. Quinn, Jr.


SUJ MAR Y


Tests were made in the Langley two-dimensional low-
turbulence tunnels of an !TiCA 653-418, a = 1.0 airfoil
section with roughness in the form of carborundtum grains
applied to the leading edge. Roughness grains having
average diameters of 0.0005 and 0.0007 airfoil chord were
applied to the leading edge of the :.ing, and lift and
drag measurements were made for a range of Reynolds num-
bers from 0.25 to 3.0 x 100. From a comparison of data
obtained in the present tests with data obtained in tests
of the smooth wing, marked reductions in maximum lift
coefficient were found to be caused by the roughness
throughout the test range of Reynolds number. The drag
coefficient at the design .ift coefficient increased
sharply and the lift-curve slope decreased rapidly at a
critical Reynolds number that depended upon the size of
the carborundum grains. This critical Reynolds number
occurred at approximately 0.50 and 0,70 x 10 for the
0.0005- and the 0.0007-chord-diameter roughness grains,
respectively. With roughness, a decrease in maximum lift
coefficient as great as 0.2, a decrease in lift-curve
slope of 0.028, and an increase in drag coefficient at
the design lift coefficient ,of C.007 were observed at a
Reynolds number of 1.0 x 106. For the smooth wing at the
same Reynolds number, the maximum lift coefficient
was 1.19, the lift-curve slope was 0.116. and the drag
coefficient was 0.0077. At Reynolds numbers greater than
1.0 x 100, the scale effect on the lift and drag charac-
teristics of the section -with both degrees of roughness
was generally in the same direction as the effect on the
lift and drag characteristics of the smooth airfoil.






2 COiFIDENTIAL NACA CB :D L5J04


INTRODUCTION


Several investigations have been made in the past to
determine the effects of Reynolds number on the aerody-
namic characteristics of various airfoil sections. A
recent investigation was made (reference 1) to determine
the effects of both Reynolds number and stream turbulence
on the lift and drag of a smooth "'.CA 6-series airfoil
section.

The present investigation was made to determine the
effects of Reynolds number on the lift and drag charac-
teristics of an NACA 6-series section with a roughened
leading edge. Tests were made, therefore, of the
NACA 653-418, a = 1.0 airfoil section in the Langley two-
dimensional low-turbulence tugnels over a range of Reynolds
number from 0.25 to 5.0 x lOb. Lift and drag measure-
ments were made at several Reynolds numbers in this range
with two degrees of roughness applied to the leading edge
of the airfoil.

Although the data presented herein and in reference 1
are quantitatively correct only for the NACA 655-418 air-
foil section, the effects of Reynolds number and roughness
would probably be in the same general direction and of
approximately the same order of magnitude for other
NACA 65-series airfoil sections that do not differ greatly
in thickness and camber from the !!-.CA 653-418. These
results are also helpful in properly evaluating the merits
of low-scale test data.


SYMBOLS


cz section lift coefficient

cd section drag coefficient

ci. design section lift coefficient

0C maximum section lift coefficient
c airfoil chorax

c airfoil chord


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NACA CB No. L5JO4


R Reynolds number

00 section angle of attack


MODELS AND TEST METHODS

The airfoil model used in the present investigation
was of 6-inch chord and was constructed of aluminum alloy
to correspond to the ordinates of the NACA 655-418, a = 1.0
airfoil section. Ordinates for this airfoil section are
presented in reference 2. A photograph of the model
is presented in figure 1.

Roughness was simulated by applying carborundum
grains of a given diameter to the leading edge of the wing
with shellac. The roughness was applied to both surfaces
of the airfoil as far back as 0.078c and the grains
covered approximately 10 percent of the roughened area.
Two degrees of roughness were obtained by use of grains
having average diameters of 0.0003c (0.002 in.)
and 0.0007c (0.003 to 0.005 in.). The standard roughness
used in systematic airfoil investigations (reference 2)
for determining the characteristics of various airfoils
having transition fixed at the nose is composed of grains
having average diameters of 0.3005c. This standard rough-
ness was thought to be considerably more severe than that
caused by any manufacturing, irregularities or poor painting
procedures but is not so severe as that caused by icing,
mud, or damage from military combat. One grain roughness
used in the present tests is larger than the standard
roughness; the other is smaller.

The tests were made in the Langley two-dimensional
low-turbulence tunnel (designated LTT) and in the Langley
two-dimensional low-turbulence pressure tunnel (desig-
nated TDT). Both tunnels are 3 feet wideand feet high and
2
were designed to test models completely spanning the jet
in two-dimensional flow. These tunnels are characterized
by air streams having exceptionally low turbulence levels,
of the order of a few hundredths of 1 percent. Lift
measurements were made by an arrangement designed to
integrate the pressure reactions along the floor and
ceiling of the tunnel test sections and drags were
measured by the wake-survey method.


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NACA CB No. L5JO4


All tests were run at tunnel Mach numbers less
than 0.2. Measurements were made at atmospheric pressure
in the LTT, and at tunnel pressures Of 1L.7, 30, 45, 63,
and 87 pounds per square inch absolute in the TDT.

Corrections for the effects of tunnel-wall interfer-
ence and air-stream constriction were applied to the
model as follows:

c, = 0.998 c't

cd = 0.-999 Cd'

where the primed quantities represent the values obtained
in the tunnel.


RESULTS AND DISCUSSION


Lift characteristics for the NACA 653-418, a = 1.0
airfoil section at various Reynolds numbers with
two degrees of leading-edge roughness are presented in
figure 2 and the drag characteristics are presented in
figure 3. In figure 2(a), in which data are presented
for the model having 0.0003c-diameter grains on the
leading edge, a pronounced jog in the lift curve is
noticeable at a lift coefficient of 0.9 at a Reynolds
number of 0.50 x 106. In reference 5 such a jog was
found to be associated with a region of laminar separa-
tion just behind the leading edge of the wing. The fact
that a jog occurred in the present tests indicates that
this degree of roughness did not completely eliminate
laminar flow at the leading edge until a Reynolds number
between 0.50 and 0.75 x 106 was attained.

From figures 2(b) and 2(c), in which data are pre-
sented for the model having 0.0007c-diameter grains on
the leading edge, no jog occurred in the lift curve at
Reynolds numbers greater than 0.35 x 106. This degree of
roughness therefore probably eliminated laminar flow ,
entirely at a Reynolds number between 0.55 and 0.50 x 106.

Curves showing the variation of maximum lift coeffi-
cient and lift-curve slope with Reynolds number are pre-
sented in figure 4 and the variation of drag coefficient
at the design lift coefficient with Reynolds number, in


COI-FPIDEITTIAL


CONFIDEI'TTAL







ITACA CB :o. L5J04


figure 5. Application of roughness to the leading edge
of the airfoil caused values of the maximum lift coeffi-
cient and the lift-curve slooe that were substantially
lower than the values for the smooth airfoil (fig. L.).
The maximum lift coefficients and the lift-curve slopes
are predominantly lower than those for the smooth wing
throughout the test range of Reynolds number, and there
is a critical Reynolds number at vhich the maximum lift
coefficient decreases noticeably and the lift-curve slope
decreases rapidly. Figure 5 shows that at this critical
Reynolds number a sharp increase in the variation of the
drag coefficient at the design lift coefficient with
Reynolds number also occurred. The lift-curve slope
decreases rapidly and the drag coefficient at the design
lift coefficient increases sharply at a Reynolds number
of approximately 0.70 x 106 for the 0.0005c-diameter grain
roughness and of approximately 0.50 x 10b for the
O.0007c-diamneter grain roughness. At a Reynolds number
of 1.0 x 100, however, the differences in values of the
lift-curve slope and of the drag coefficient for the
two degrees of roughness disappear, and at greater
Reynolds numbers, within the accuracy of the results, the
values of these quantities appear to be independent of
the sizes of the roughness for which data are presented.

The lift-curve slopes in figure h also show that for
the 0.0003c-diameter grain roughness the lift-curve slope
is essentially the same as for the smooth wing up to a
Reynolds number of at least 0.50 x 10b. This degree of
roughness probably brought about n.o significant changes
at low lift coefficients in the development of the
boundary layer from that existing on the smooth wing up
to a Reynolds number of 0.50 x 106. Because the maximum
lift coefficient for this degree of roughness ,vas lower
than that for the smooth wing throughout the entire range
of test Reynolds numbers, the roughness probably did
induce some change in the nature of the flow at high lift
coefficients.

With the 0.0007c-diameter roughness grains, the lift-
curve slope and maximum lift coefficient were greater than
those of the smooth wing at a Reynolds number of 0.25 x 106,
but at Reynolds numbers greater than 0.30 x 106 these
quantities were lower than those of the smooth airfoil
for both degrees of roughness. The reason for this phe-
nomenon is not readily evident. There is a possibility,
however, that at a Reynolds number of 0.25 x o106 the
roughness was not large enough to destroy the laminar


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NACA CB No. L5JOi


flow entirely but was large enough to prevent laminar
separation behind the minimum pressure point. A boundary-
layer velocity distribution would result, therefore, which
would be different from both the smooth flow condition
and the smaller roughness condition.

Figure shows that at Reynolds numbers between 1.0
and 3.0 x 10O the maximum lift coefficients were less than
those of the smooth wing by. approximately 0.14 and 0.20
for the 0.0003c- and 0.0007c-diameter roughness grains,
respectively. At a Reynolds number of 1.0 x 106, the
lift-curve slope of the rough wings was 24. percent less
than that of the smooth wing, but at a Reynolds number
of 3.0 x 106, a decrease due to roughness of approximately
12 percent in lift-curve slope was found. A constant
increment in drag coefficient at the design lift coeffi-
cient due to roughness of approximately 100 percent was
found (fig. 5) at Reynolds numbers between 1.0 and 3.0 x 106.


CONCLUSIONS


A comparison of results of tests of the
NACA 655-148, a = 1.0 airfoil section for a range of
Reynolds number from 0.23 to 3.0 x 106 with roughness
grains having average diameters of 0.0003 and 0.0007 air-
foil chord (0.0003c and 0.0007c) with results of previous
tests of the smooth wing led to the following conclusions:

1. Maximum lift coefficients of the airfoil with
roughness were generally lower than those obtained on the
smooth airfoil section throughout the test Reynolds number
range. At a Reynolds number of 1.0 x 106 the maximum lift
coefficient for the smooth wing was reduced from a value
of 1.19 to 1.05 and 0.99 by the 0.0005c- and
0.0007c-diameter grains, respectively.

2. There is a critical Reynolds number at which the
lift-curve slope decreases rapidly and the drag coeffi-
cient increases sharply depending upon the size of the
roughness. This critical Reynolds number was approxi-
mately 0.70 and 0.50 x 106 for the 0.00053c- and
0.0007c-diameter grains, respectively.

3. With roughness, at a Reynolds number of 1.0 x 106,
the lift-curve slope was 0.088 and the drag coefficient
at the design lift coefficient was 0.0155 whereas the


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NACA CB No. L5Jok


corresponding values for the smooth airfoil section
were 0.116 and 0.0077, respectively. At Reynolds numbers
greater than 1.0 x 10o the changes in lift-curve slope
and Crag coefficient were nearly independent of the sizes
of the roughness for the two degrees of roughness for
which the effects 'were measured.

4. Large variations in the lift and drag character-
istics of the airfoil were found in the range of Reynolds
number between 0.25 and 1.0 x 106. At Reynolds numbers
greater than 1.0 x 10', the scale effect on the lift and
drag characteristics of the section with both degrees of
roughness was generally in the sane direction as the scale
effect on the characteristics :f the smooth airfoil.


Langley Memorial Aeronautical Laboratory
National Advisory Committee for Aeronautics
Langley Field, Va.





REFERENCES


1. Quinn, John H., Jr., and Tucker, barren A.: Scale and
Turbulence Effects on the Lift and Drag Character-
istics of the IIACA 65 3-41, a = 1.0 Airfoil Section.
NACA ACR No. L4-H11, 1J44.

2. Abbott, Ira H., von DocEnhoff, Albert E., and Stivers,
Louis S., Jr.: Surimary of Airfoil Data. NACA ACR
No. L5C05, 1945-

3. von Doenhoff, Albert E., and Tetervin, Neal: Investi-
gation of the Variation of Lift Coefficient with
Reynolds Number at a Moderate Angle of Attack on a
Low-Drag Airfoil. IACA CB, Nov. 1942.


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