The relation between spanwise variations in the critical mach number and spanwise load distributions

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Material Information

Title:
The relation between spanwise variations in the critical mach number and spanwise load distributions
Series Title:
NACA WR
Alternate Title:
NACA wartime reports
Physical Description:
5 p., 3 leaves : ill. ; 28 cm.
Language:
English
Creator:
Whitcomb, Richard T
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:
Airplanes -- Wings -- Testing   ( lcsh )
Aerodynamics -- Research   ( lcsh )
Genre:
federal government publication   ( marcgt )
bibliography   ( marcgt )
technical report   ( marcgt )
non-fiction   ( marcgt )

Notes

Summary:
Summary: Data are presented to show the changes that occur in the spanwise load distributions on wings when the critical Mach number is exceeded. These data indicate that the magnitude of the changes in spanwise load distribution varies with the magnitude of the spanwise variation in the critical Mach numbers of the sections. Means of reducing the magnitudes of such changes are considered.
Bibliography:
Includes bibliographic references (p. 5).
Statement of Responsibility:
by Richard T. Whitcomb.
General Note:
"Originally issued December 1944 as Advance Confidential Bulletin L4L07."
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 - 003806637
oclc - 124093586
System ID:
AA00009434:00001

Full Text


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7E 1BLATION BETWEEN SPANWISE VARIATIONS IN THE CRITICAL

MACH NUMBER AND SPANWISE LOAD DISTRIBUTIONS


By Richard T. Whitcomb


Langley Memorial Aeronautical Laboratory
Langley Field, Va.

UNIVERSITY OF FLORIDA
DOCUMENTS DEPARTMENT
120 MARSTON SCIENCE LIBRARY
P.O. BOX 117011
GAINESVILLE, FL 32611-7011 USA


WASHINGTON


SNACA 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-
i:,vitly-held under a security status but are now unclassified. Some of these reports were not tech-
.: ally edited. All have been reproduced without change in order to expedite general distribution.


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NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS







WAIRTI'IIME RE PORT

ORIGINALLY ISSUED
December 1944 as
Confidential Bulletin L4L07


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NACA CB No. L4L07

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS


CONFIDENTIAL BULLETIN


THE RELATION BETWEEN SPANWISE VARIATIONS IN THE CRITICAL

MACH NUMBER AND SPANNISE LOAD DISTRIBUTIONS

By Richard T. Whitcomb


SU.MMAR Y


Data are presented to show the changes that occur in
the spanwise load distributions on wings when the critical
Mach number is exceeded. These data indicate that the
magnitude of the changes in spanwise load distribution
varies with the magnitude of the spanwise variation in
the critical Mach numbers of the sections. Means of
reducing the magnitudes of such changes are considered.


INTRODUCTION


The results of tests of numerous airfoils at high
speeds indicate that there may be considerable changes
in the spanwise load distributions on a wing when the
critical Mach number of the wing is exceeded. After
the local Mach number on an airfoil section exceeds a
value of unity, a compression shock is formed that
results in a decrease in the lift coefficient on the
section for the same angle of attack (references 1 to 3).
These decreases of lift coefficient generally occur at
different flight speeds on the various sections of a
wing. A change in the spanwise load distribution would
therefore usually be expected to occur on a wing
operating at a Mach number above that at which a loss
of section lift coefficient first occurs since, at
this Mach number, some sections will have experienced
a greater loss in lift than other sections. Such changes
affect the wing bending moments, the airplane trim, and
the stability characteristics.

The possibility that such changes may occur has
been recognized for several years (reference 4). A
means is available for estimating the magnitude of
the changes through the use of the low-speed lifting-
line theory and two-dimensional high-speed wind-tunnel
data (reference 5).








NACA CB No. L4LO7


An analysis of the results of wind-tunnel tests
in two-dimensional flow (references 1 to 3) indicates
that the magnitude of the loss in lift coefficient which
occurs at supercritical Mach numbers is a function of
the amount by which the operating Mach number exceeds
the critical Mach number. The magnitudes of the changes
in the spanwise distribution of load on most airplane
wings should therefore be expected to vary with the
magnitude of the spanwise variations of the section
critical Mach number.

The purpose of the present paper is to illustrate
the relationship between the spanwise variations of
load distribution and the section critical Mach number.
In order to show this relationship, subcritical and
supercritical load distributions and variations of
critical Mach number, calculated from pressure measure-
ments made during high-speed wind-tunnel tests, are
shown for three tapered wings. Means of reducing the
indicated changes are considered.


EXPERIMENTAL RESULTS


Figure 1 shows the saanwise load distributions for
a wing (NACA 23015 root section and NACA 4412 tip
section) on which there is a large spanwise variation
in the section critical Mach number. Figure 2 shows the
spanwise load distributions for a wing (Boeing 117 sec-
tion, 22 percent thick at root and 12 percent thick at
tip) on which the spanwise variation in the section
critical Mach number is moderate. These two wings were
tested in the Langley 8-foot high-speed tunnel. Figure 5
shows the snanwise load distributions for a wing
(NACA 63('l2b.-.22 root section and NACA 63(420)-517 tip
section) on which there is only a slight spanwise
variation in the section critical Mach number. This
wing was tested in the Ames 16-foot high-speed tunnel.
The distributions have been determined for a wing lift
coefficient of 0.2. The suDercritical span loadings
are for the Mach numbers at which the variations from
the subcritical loadings are most pronounced. The
loadings are presented in the conventional manner -
that is, as cnc plotted against the distance from
the center of the wing along the semispan, where Cn is
the section normal-force coefficient and c is the
section chord. The spanwise variations in the section


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NACA C0 No. L L07


critical Mach number Mcr are also shown in the figures.
These variations in critical Mach numr.ber are determined
for the angles of attac!c corrasyon.iing to a wing lift
coefficient CL of 0.2 at low speeds.

The results shown were obtained from data recorded
during tests of winrr models that spanned the throats of
the tunne's. The air flow over the wing sections near
the tips was t'.erefore aonroxitm tely two dimensional as
compared with the flow that woijld have been present
had the wing been testpl with free tips. Tinpiblished
data obtained during wind-tunnel tests at high speeds
of a wing with a fr-e tip indiceLe that the critical
Mach number of a wing section is r.-a-ter when the
section is operating in the air flo.J near a free tip
than when the section is operatir.e .'n a two-dinensional
flow. If the wins had been tes:.eJ with free tips at
the same angles of attack, the critical Mach numbers
of the tip sections would therefore probably have been
greater than the critical M'ach numbers measured during
the present tests.





The results pre., nted in f i ure 1 show the radical
changes that can occur in soanwise load distribution on
a wing with a large saanwise variation in the section
critical Mach number. A cnnsiderable clancre in load
distribution on a winc- with a moderate sno.nwise variation
in the critical .ac 1-.m i-ber is shown in figure 2, and
a negligible chinge on a w.'ing ,ith a. small soanwise
variation in the critical rlach number is shlcwn in
figure 5. These exoeri.mental results thus indicate
that the m-agnitude of the rhan-e in the snanwise load
distribution on a win" at supercritical Mach numbers
varies with th' .-ra-niti_'d: of th, spanwise variation
in the critical hlach n-mbc er of the wir,; sections.

The outboard mov.-ment of c!.E c -nLer ofi load on the
semispan of a 'ving at supercritical &;,ach numbers, as
shown in figure-s 1 arn.- 2, d-cre: ses t',e lownwash in the
region of the tail for a giv-n ,-, lift c,--officient
and decreases the change in down'wash "or a given change
in wing lift coefficient. These variations change the
elevator deflection required for trim and increase the


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NACA CB No. L L07


stability of the airplane. Such an outboard movement
of the center of load also increases the bending
moments on a winj structure. If the outboard movement
occurs when the wing is supporting its maximum design
load, the factor of safety for the wing structure is
decreased. The decrease in the lift coefficient on a
wing for a given angle of attack at supercritical Mach
numbers requires that the angle of attack of the air-
plane be increased in order to maintain a given lift
coefficient. This increase in angle of attack leads to
changes in the elevator deflection required for trim and
to increases in the stability at supercritical Mach
numbers, in addition to the changes produced by a
spanwise movement of the center of load.

Because the outboard movement of the center of
load produces detrimental changes, this movement should
be held to a minimum. A comparison of results in
figures 1 to 3 indicates that, for a definite moderate
lift coefficient, the outboard movement can be reduced
by designing the wing-fuselage combination to give the
same critical Mach number for each of the wing sections.
The obvious method of obtaining this result is to design
the wing with the same section and the same section lift
coefficient at each station and to reduce to a minimum
the interference effects on the wing. The results
presented in figure 5 indicate that the same result may
be accomplished by using the proper combination of
various wing sections. The outboard movement of the
center of load may also be reduced by so deflecting
"dive-recover-" flaps placed inboard on the lower
surfaces of the wing that the lift increases on the
inboard sections where the greater losses in lift occur.
Dive-recover", flaos placed outboard would increase
rather than decrease the wing bending moments for a
given lift and would be less effective than inboard
flans in reducing the total changes in the elevator
deflection required for trim and in reducing the
stability of an airplane.


CONCLUDING REMARKS


A comparison of the results of tests of three
different tapered wings indicates that the magnitude
of the spanwise movement of wing center of load at
supercritical Mach numbers varies with the magnitude
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NACA CB No. L)EL07 CONFIDENTIAL 5

of sDanwise variation in critical Mach number; conse-
quently, it may prove desirable in the design of wings
for high-speed airplanes to choose sections, thickness-
to-chord ratios of sections, and section load distri-
butions to provide a constant value of the spanwise
critical Mach number.

Some of the effects of the spanwise shift of the
center of load may be overcome by the use of dive-
recovery flaps placed inboard.


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



PREFERENCES

1. Stack, John, Lindsey, '.V. F., and Littell, Robert E.:
The Compressibility Burble and the Effect of
Compressibility on Pressures and Forces Acting
on an Airfoil. NACA Rep. No. 646, 1958.

2. Stack, John, and von Doenhoff, Albert E.: Tests of
16 Related Airfoils at High Speeds. NACA
Rep. No. 492, 1954.

5. Stack, John: The N.A.C.A. High-Speed Wind Tunnel
and Tests of Six Propeller Sections. NACA Rep.
No. 463, 1955.

4. Sibert, H. W., and Lees, Lester: Compressibility
Phenomena as Related to Airplane Structural Design.
ACT.R No. 4524 (Revision I), Material Div.,
Air Corps, July 7, 1942.

5. Boshar, John: The Determination of Span Load
Distribution at High Speeds by Use of High-Speed
Wind-Tunnel Section Data. NACA ACR No. 4B22, 1944.


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NACA CB No. L4L07


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Figure 1.- Measured spanwise load distributions and apanwise
variation of section critical Mach number on tapered wing
with NACA 23015 root section and NACA 1412 tip section at
CL = 0.2.


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NACA CB No. L4LO?7


Tunnel sw/e


Front view of
wing


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NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS


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Figure 2.- Measured spanwise load distributions and spanwise
variation of section critical Mach number on tapered wing
with root-section thickness ratio of 22 percent and tip-
section thickness ratio of 12 percent at CL = 0.2.


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Front view of
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'emir'span CON FIDENTIAL
Figure 3.- Measured spanwise load distributions and spanwise
variation of section critical Mach number on tapered wing
with NACA 63(420)-k22 root section and NACA 65(M20)-517
tip section at CL = 0.2.


CONF IDENTICAL


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UNIVERSITY OF FLORIDA

3 1262 08106 55 2



UNIVERSITY OF FLORIDA
-,-CUMENTS DEPARTMENT
.0. j MARSTON SCIENCE UBRARY
.- BOX 117011
.4ES\ILLE, FL 32611-7011 USA


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