The flow past a straight- and a swept-wing-body combination and their equivalent bodies of revolution at Mach numbers ne...

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Material Information

Title:
The flow past a straight- and a swept-wing-body combination and their equivalent bodies of revolution at Mach numbers near 1.0
Series Title:
NACA RM
Physical Description:
18 p. : ill. ; 28 cm.
Language:
English
Creator:
Lindsey, W. F
Langley Aeronautical Laboratory
United States -- National Advisory Committee for Aeronautics
Publisher:
NACA
Place of Publication:
Washington, D.C
Publication Date:

Subjects

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

Notes

Abstract:
Abstract: The complete flow fields past a straight and a swept wing-body combination and their equivalent bodies of revolution at Mach numbers around 1.0 have been observed by means of motion-picture schlieren photography. The results of these observations indicate that the shock growth and positions on the wing-body combinations are closely reproduced in the flow past their respective equivalent bodies.
Bibliography:
Includes bibliographic references (p. 6).
Statement of Responsibility:
by Walter F. Lindsey.
General Note:
"Report date January 13, 1954."
General Note:
"Classification changed to unclassified Authority: NACA Research Abstract No. 96 February 10, 1956."--stamped on cover

Record Information

Source Institution:
University of Florida
Rights Management:
All applicable rights reserved by the source institution and holding location.
Resource Identifier:
aleph - 003808476
oclc - 129537481
sobekcm - AA00006156_00001
System ID:
AA00006156:00001

Full Text


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LANDUM


'LOW PAST A STRAIGHT- AND A SWEPT-WING-BODY COMBINATION
AND THEIR EQUIVALENT BODIES OF REVOLUTION

AT MACH NUMBERS NEAR 1.0

By Walter F. Lindsey
Langley Aeronautical Laboratory
Langley Field, Va.
SSTY OPF FROO CIASSFICATION CHANGED TO UNCLA LJFID
p MARSTON SCIENCE LIBRARY AUTERT: NACA REEARC ABSTRACT NO. 9
X 117011
E, FL i32611-7011 USA DATE: FEBR;UAR 10, 1956 v.
c "SUIFIED DOCUMENT;
~;,1Ye ~M n metal amltarstim tectliu um Ibtuaml Deleai of the United Sltees mLtUn the meanlng
of the FnIarmw bun, TIek 1, U..C., teim. 7W am 1M, uthe riamnnion or reielationu oF which In any
oenar in. a mm nse pasa. s proablhied lr InE.

CIATIONAL ADVISORY COMMITTEE
... FOR AERONAUTICS
WASHINGTON
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NACA RM L54A28a CONFIDENTIAL

NATIONAL ADVISORY COMM -ITTEE FOR AERONAUTICS


RESEARCH MEMORANDUM


THE FLOW PAST A STRAIGHT- AND A SWEPT-WING-BODY COMEI.IATIIr

AND THEIR E ,U IVALENT BODIES OF REVOLUTION

AT MACH NUMBERS NEAR 1.0

By Walter F. Lindsey


SUMMARY


Tests utilizing the schlieren method of flow photography have been
conducted to provide a comparison of the complete flow fields past a
straight- and a swept-wing-body combination and the flow fields past
their equivalent bodies of revolution at Mach numbers around 1.0. The
results indicated that the shock growth and positions on the wing-body
combinations were closely reproduced in the flow past the equivalent
body.


INTRODUCTION


The correspondence of drag between a wing-body combination and its
equivalent body of revolution as determined by the transonic "area rule"
was stated in reference 1 to require a correspondence of the flow fields
and shock formation about the two bodies. Some limited information on
the flow fields in reference 1 indicated that such a condition existed.
In the absence of information whereby a comparison of the complete flow
fields past a wing-body combination and its equivalent body could be
made, tests utilizing the schlieren method of flow photography have been
conducted to provide this information by examining the complete flow
fields and the shock formations existing on two wing-body combinations
and their equivalent bodies.


APPARATUS AND TESTS


The tests were conducted in the Langley 4- by 19-inch semiopen
tunnel (ref. 2), which had been modified to operate on a direct-blowdown
principle. A support sting whose diameter was equal to the body diameter


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at the body-sting juncture and increased downstream (fig. 1) was
installed in the tunnel. All models involved in this investigation
were supported at an angle of attack of 00 on the sting.

The models consisted of body alone, two wing-body combinations,
and their equivalent bodies. The basic body used in all combinations
had a fineness ratio of 12 and a profile as shown in figure l(a). The
wing of one combination was a 450 sweptback wing having a taper ratio
of 1 and an aspect ratio of 3.9. The wing had an NACA 66-006 airfoil
section in a streamwise direction. The equivalent body had the same
longitudinal area distribution and is shown in figure 1. The wing of
the second wing-body combination was a straight wing having a taper
ratio of 1 and an aspect ratio of 1.5. The wing had a 6-percent-thick
symmetrical circular-arc profile with maximum thickness at 50 percent
chord. The profile and area distribution of the equivalent body for
this combination are also shown in figure 1. The profiles of the
equivalent bodies for both wing-body combinations can be compared in
figure l(a). The bump on the basic body formed by the addition of the
cross-sectional area of the swept wing, in accordance with the transonic
area rule (ref. 1), extended along the body for a distance of 40 percent
of the body length and had a maximum height (increment in radius) of
17 percent of the basic body radius. The bump formed by the addition
of the cross-sectional area of the straight wing extended for a distance
of 25 percent of the body length and had a maximum height of 41 percent
of the body radius. The bump for the swept wing thus represents a
relatively small disturbance on the body, whereas, the bump for the
straight wing represents a moderately large disturbance.

Data on the flow fields of the models were obtained on 35-mm film
in the form of schlieren photographs taken as motion pictures of the
flow; these data were obtained for a slowly but continuously increasing
Mach number over the speed range, followed by a similar decrease in
Mach number. The variable-frequency light source described in refer-
ence 3 was used and limited each picture to an exposure of about
4 microseconds. The Mach number for any given photograph was identified
by a Mach number indicator that extended into the lower part of the field
of observation. The bodies, as previously stated, were mounted at an
angle of attack of 00 and the wings on the body were oriented with
respect to the optical axis of the schlieren system at 90, 450, and 00
For the 00 view, the wing was along the optical axis of the schlieren
system and was not observable in the schlieren pictures.

The tests were conducted over a Mach number range from approximately
0.7 to slightly above 1.0 at a constant st.jg7nation pressure of 20 psia.
The corresponding Reynolds number based on body length at a Mach number
of 1 was 3 x 106.


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RESULTS AND DISCUSSION


Selected photographs from the motion pictures are presented for
the body alone and for the wing-body combinations at Mach numbers of
0.95, 0.98, 1.0, and approximately 1.04, since these photographs cover
the Mach number range of shock development. Photographs of the flow
past the body alone (fig. 2) show that the body shock starts formring
around a Mach number of 0.98 and moves rearward along the body with
increasing Mach number. The addition of carborundum on the nose of the
body to fix transition of the boundary layer had no effect on the shock
pattern nor on the location of the shocks except to produce character-
istic disturbance patterns at the roughness location. (Compare figs. 2
and 5.) The similarity of shock formation indicates that the turbulent
boundary layer produced by carborundum is not affecting the shock forma-
tion and that a smooth model may be used thrc_-,hout the tests without
any appreciable effect of sudden transition in the boundary layer on
the smooth-nose models.

In order to examine carefully the flow about the vin;:-b-dy com-
binations, three views of the flow were made one taken normal to the
span of the wing (900 view), the second taken along the span of the
wing (0 view), and the third taken midway between these two (45 view).
From an examination of the flow in the three views, the axially symmetri-
cal or dissymmetrical nature of the flow past the wing-body combinations
can be determined. The three views of the flow past the 45 swept-wing-
body combination and the views of the flow past its equivalent body at
Mach numbers from 0.95 to about 1.04 are shown in figure 4. At a Mach
number of 0.95 (fig. 4(a)) only minor shocks are observed in the three
views and there is close correspondence of the flow past the wing-body
and the equivalent body. An increase in the Mach number to O.,9'
(fig. 4(b)) produces somewhat stronger shocks; and, again, similar
flow conditions are observed on the wing-body and equivalent body. At
this Mach number the body shock is formed ahead of the wing. At a
Mach number of 1.0 (fig. 4(c)), the shock has moved downstream and is
located near the wing-tip trailing edge and extends completely across
the body. The three views of the flow past the wing-body combination
indicate an approximately axially symmetrical type of flow, which is
well-reproduced by the equivalent body not only in aIpe-rmnce of shock
strength but also in the locations of the body shock ahead of the wing
and the main shock. At a Mach number of 1.04, there is some variation
in the shock appearance around the body in the wing-body combination or
some axial dissymmetry of the shock. This variation is slight, however,
and the shock pattern is well-duplicated on the equivalent body. The
flow conditions observed past the wing-body combination and its equiva-
lent body indicate that throughout the speed range of the tests the
appearance and growth of the shocks on the wing-body combination are
well-duplicated on the equivalent body, whose distortion from the basic


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NACA RM L54A28a


body represents a small disturbance. Although this result indicates
equivalence of drag on an untapered sweptback wing-body combination and
its equivalent body, some investigations refss. 1, 4, and 5) of t iered
sweptback wing-body combinations and their equivalent bodies have shown
the increment in drag at transonic speeds for the equivalent body to
be between 40 and 80 percent of that for the winr-body combination.

The flow past the str. ht-wing-body combination and its equivalent
body is shown in figure 5. At a Mach number of 0.95 (fig. 5(a)), the
flow past the wing-body combination ;.r-e o's similar in the different
views; thus, approximately axial symmetry of shock formation is indi-
cated. The shock location somewhat ahead of the wlin trailing edge is
well-duplicated on the equivalent body. At a Mach number of 0.98
(fig. 5(b)), the shock is at the trailing edge of the wing and little
difference is noted in the three different views of the flow past the
body. The shock location is approximately the same as on the equivalent
body, yet the equivalent body does produce a small difference in Lhe
shock at the shock-body juncture where the shock is forked. At a Mach
number of 1.0 (fig. 5(c)), the body shock ahead of the wing is occurring
near the same location in the three views of the flow past the wing-body
combination as on the equivalent odj;.. The main shock behind the win,;,
however, does show, in the three different views, considerable devia-
tion from axial symmetry. The flow past the equivalent body has the
same location of terminal shock and body shock and produces a terminal
shock whose shape thr-u,_holt the field appears to be an average of the
variously shaped terminal shocks observed on the wing body. The flow
past the equivalent body is an axially symmetrical representation of
the axially dissymmetrical flow past the wing-body combination. A
similar condition is observed at a Mach number of approximately 1.04
(fig. 5(d)).

These results indicate that, when the winr-bo'--;dy combination gives
rise to rapid changes in area along the axis of its equivalent body and
represents a moderately large disturbance, axial dissymmetry of the
shock formation can occur, yet the equivalent body provides a flow con-
dition that closely corresponds to an integrated effect of the dis-
symmetry in the aforementioned shock formation. The flow past the
equivalent body would be expected to provide increments in dra, due to
shock that would closely -:rresLond to the increments in drag due to
shock on the wing-body combination.

The results of the tests on the two wing-body combinations have
indicated that the flow past the equivalent body duplicates or closely
m:ppr'.:irm.es that past the wirn:-body combinations; furthermore, the
results indicate that, as the body sh'.pe, or the bump representing the
equivalent body, departs radic ally from a slender bump or small dis-
turbance, the similarity of flow between the wing-body combination and
its t(uivalent body divr-i-s. Presumably, therefore, a wing-body


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NACA RM L54A28a CONFIDENTIAL 5


combination having a straight wing of aspect ratio greater than 1.5 and
thickness-chord ratio greater than 6 percent would not only have con-
siderable axial dissymmetry of the flow past the model, but the differ-
ences between the flow past the model and the flow on its equivalent
body would be larger than those shown herein.


Langley Aeronautical Laboratory,
National Advisory Committee for Aeronautics,
Langley Field, Va., January 15, 1954.


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F T-T F'E i 'iES


1. Whitcomb, Richard T.: A Study of the Zero-Lift Drag-Rise Character-
istics of Wing-Body Combinations Near the Speed of Sound. NACA
RM L52H08, 1952.

2. Daley, Bernard N., and Dick, Richard S.: Effect of Thickness, Camber,
and Thickness Distribution on Airfoil Characteristics at Mach Num-
bers Up to 1.0. NACA RM L52G51a, 1952.

5. Lindsey, Walter F., and Burlock, Joseph: A Variable-Fr- ji.r- n:y Light
Synchronized With a High-Speed Motion-Picture Camera To Provide
Very Short Exposure Times. NACA TN 2949, 1955.

4. Hoffman, Sherwood: An Investigation of the Transonic Area Rule by
Flight Tests of a Sweptback Wing on a Cylindrical Body With and
Without Body Indentation Between Mach Numbers 0.9 and 1.8. NACA
RM L53J20a, 1955.

5. Hall, James Rudyard: Comparison of Free-Flight Measurements of the
Zero-Lift Drag Rise of Six Airplane Configurations and Their
Equivalent Bodies of Revolution at Transonic Speeds. NACA
RM L55J21a, 1955.


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DOCUMENTS DEPARTMENT
120 MARSTON SCIENCE LBRARY
RO. BOX 117011
GAINESVILLE, FL 32611-7011 USA


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