A study of the flow over a 45° sweptback wing -fuselage combination at transonic Mach numbers

MISSING IMAGE

Material Information

Title:
A study of the flow over a 45° sweptback wing -fuselage combination at transonic Mach numbers
Series Title:
NACA RM
Physical Description:
60 p. : ill. ; 28 cm.
Language:
English
Creator:
Whitcomb, Richard T
Kelly, Thomas C
Langley Aeronautical Laboratory
United States -- National Advisory Committee for Aeronautics
Publisher:
NACA
Place of Publication:
Washington, D.C
Publication Date:

Subjects

Subjects / Keywords:
Aerodynamics, Transonic   ( lcsh )
Airplanes -- Wings, Swept-back   ( lcsh )
Aerodynamics -- Research   ( lcsh )
Genre:
federal government publication   ( marcgt )
bibliography   ( marcgt )
technical report   ( marcgt )
non-fiction   ( marcgt )

Notes

Abstract:
Abstract: Pressure distributions, tuft patterns, and schlieren surveys have been obtained for a 45° sweptback wing-fuselage combination in the Langley 8-foot transonic tunnel at transonic Mach numbers to 1.11 and angles of attack to 20°. The results provide an indication at transonic Mach numbers of the nature of the formation of shock waves on the wing and fuselage, wing-fuselage interference, and the development of separation and the separation vortex.
Bibliography:
Includes bibliographic references (p. 14).
Statement of Responsibility:
by Richard T. Whitcomb and Thomas C. Kelly.
General Note:
"Report date June 25, 1952."

Record Information

Source Institution:
University of Florida
Rights Management:
All applicable rights reserved by the source institution and holding location.
Resource Identifier:
aleph - 003809146
oclc - 133095767
sobekcm - AA00006202_00001
System ID:
AA00006202:00001

Full Text
... ... .
,- P F :r:..
iII .. ,.c ". ". ... '- '


/RM
RM L52D01


RMLSDO


SS .... ;:;,. : .:. *; .







CH MEMORANDUM









'.'.. /By Richard T. Whitcomb and Thomas C. Kelly

::':.;angley Aeronautical Laboratory
i: : ':.'. Langley Field, Va.



UNVERs OF FLORIDA
DlOCUMENTS DEPARTMENT
i -O. BOX 11.701 1






FOR. AERONAUTICS
S.:6......;. :: WASHINGTON ,
I JIe 25I. 952
". 95
[ "i" '! 'u. 25,"
""1.5" t{''" "
) .&M.: ,'% .,. .:.,: ,.


June 25~ 1952


-:, : ,.. ...., : ,:.. :'. ,,..:,.. : .. ,.
i
A lli, 4 4.. ,^. T ...
_ ,,',' ;, :. .. .: ; : .


I







.i









































































































:'





13 ?

NACA RM L52D01

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

RESEARCH MEMORANDUM


A STUDY OF THE FLOW OVER A 450 SWEPTBACK WING-FUSELAGE

COMBINATION AT TRANSONIC MACH NUMBERS

By Richard T. Whitcomb and Thomas C. Kelly


SUMMARY


Pressure distributions, tuft patterns, and schlieren surveys have
been obtained for a sweptback wing-fuselage combination in the Langley
8-foot transonic tunnel at Mach numbers to 1.11 and angles of attack
to 200. The wing had 450 of sweepback, an aspect ratio of 4.0, a taper
ratio of 0.6, and an NACA 65A006 airfoil section. A study of the results
of these measurements indicates the development of various phenomena with
increases in Mach number and angle of attack. Among these phenomena are
the development of the shock on the wing, the initiation and rearward
movement of a strong normal shock behind the trailing edge of the wing-
fuselage juncture, the onset of the bow shock ahead of the wing leading
edge, and the increase and reduction of the boundary-layer separation
and the leading-edge boundary-layer vortex.


INTRODUCTION


Several detailed wind-tunnel investigations refss. liand 2, for
example) have provided a basis for the understanding of the flow over
sweptback wings at high-subsonic Mach numbers. On the basis of these
data and pressure data obtained from the wing-flow method (ref. 3) the
nature of the flow over sweptback wings at transonic speeds has been
conjectured. Because of the previous speed limitations of wind tunnels,
however, it has been impossible to obtain a more detailed investigation
of the nature of this flow at transonic speeds. A slotted test section
which allows an investigation of relatively large models in the transonic
range to a Mach number of 1.14 recently has been installed in the Langley
8-foot transonic tunnel. With this new facility, a detailed investigation
of the flow phenomena over a 450 swept-wing fuselage combination has
been made. The results of this study provide not only a contribution to
the knowledge of the flow over swept wings in the transonic range but
also an indication of the nature of sweptback wing-fuselage interference
at transonic Mach numbers.







NACA RM L52D01


The data to be discussed include pressure distributions, tuft
patterns, and schlieren surveys. Through consideration of these data
it has been possible to present a qualitative description of development
of shock waves and boundary-layer separation on the wing and fuselage at
transonic Mach numbers.


APPARATUS


The Langley 8-foot transonic tunnel is a single-return, dodecagonal,
slotted-throat wind tunnel which operates at a stagnation pressure
approximately equal to atmospheric pressure. The tunnel is capable of
continuous operation up to a Mach number of 1.14. A complete description
of the Langley 8-foot transonic tunnel may be found in reference 4.

The model configuration for the present investigation had a wing
with 45 sweepback of the quarter-chord line, an aspect ratio of 4, a
taper ratio of 0.6, and an NACA 65A006 airfoil section parallel to the
air stream. The fuselage of the combination, which is shown in figures 1
and 2, had a fineness ratio of 10 based on model diameter and length from
the model nose to intersection with the sting. Two models were used to
obtain these data. One, used for pressure measurements, had a wing con-
structed of a mild steel core with a tin-bismuth-alloy covering and is
described in reference 5; the other, used for schlieren surveys, tuft
surveys, and force measurements, had a wing constructed of 14S-T aluminum
alloy and is described in reference 6. The model used for pressure
measurements is shown mounted in the 8-foot slotted test section in
figure 1. General dimensions of the models and locations of pressure
orifices on the wing and fuselage are presented in figure 2.

Turt surveys were made with alternate rows of nylon line and wool
yarn cemented directly to the surface of the model. The very flexible
wool-yarn tufts gave a good indication of slight changes in flow direc-
tion. The less flexible thin nylon tufts remained on the wing longer
at higher Mach numbers, however, and gave a good indication of violent
separation. It was found also that the difference in the thicknesses
of the two types of tufts could be used to determine the relative thick-
ness of the boundary layer. Schlieren photographs were made with the
temporary single-pass system described in reference 4.


RESULTS


The data to be analyzed are presented as individual groups for given
survey conditions. Each gro0p consists of pressure data, force data, tuft







NACA RM L52D01 3


patterns, and schlieren surveys. Survey conditions for various Mach num-
bers at each of several angles of attack are presented in figures 3 to 8.

Pressure data to be analyzed are presented in the form of pressure-
coefficient profiles plotted at the five semispan measurement stations
on a plan form of the wing and at six radial locations on an outline of
PZ PO
the fuselage. The pressure coefficient P is defined as
qo
where p and qu are the free-stream static and dynamic pressures,
respectively, and p2 is the local static pressure. The pressure data
were taken directly from the tabulated data of references 5 and 7. Force
data anid tuft-pattern photographs are presented with each pressure-
coefficient profile. Variations of the force results with Mach number
are presented in figure 9. Force and tuft data were taken from more
complete unpublished data obtained in the Langley 8-foot transonic tunnel.

Schlieren data are presented for most cases in the form of composite
side views and plan views of the model. The plan-view composite is placed
at the correct spanwise location for 00 angle of attack. The axial loca-
tionS of the wing root and tip are also shown in the schlieren side view.
Photographs used to construct the composite side view were obtained by
using a stationary schlieren system and moving the model both longitudi-
nally and vertically in the test section. The plan views were obtained
by rotating the model 900 and offsetting it vertically. Because the
individual pictures used in the schlieren composites were taken during
separate runs, slight variations in the tunnel Mach number result in
discontinuities of the various shocks as they extend from one picture
of the composite to another. The grid lines shown in most of the
schlieren photographs are approximately parallel and normal to the flow.
The object shown above the rearward end of the fuselage for the 40 angle-
of-attack case is a probe which was used to measure fluctuations in down-
wash angle. The probe had no noticeable effect on the schlieren indica-
tions. It should be noted that the scale of the schlieren composites and
that of the pressure profiles are not equal.


DISCUSSION


For convenience, the discussion is divided into considerations of
the phenomena at individual angles of attack and Mach numbers. Through-
out these individual discussions, however, the development of various
phenomena with increases in Mach number and angle of attack will be noted.
Among these phenomena are the expansion of the field of flow of the body,
the development of the shock on the wing, the initiation and rearward







NACA RM L52D01


movement of a strong normal shock behind the trailing edge of the wing-
fuselage juncture, the onset of bow shock ahead of the wing leading edge,
and the increase and reduction of the boundary-layer separation and the
leading-edge boundary-layer vortex.


Angle of Attack of 0

At an angle of attack of 0 and a Mach number of 0.85 (fig. 3(a))
the pressure distributions and tuft surveys indicate the presence of
typical subcritical flow on the wing and fuselage. The drag is similar
to that at other subsonic Mach numbers and disturbances in the field
about the model are slight, as indicated by the schlieren photographs.
The pressure data of reference 5 for the fuselage-alone configuration
indicate that the increase in the velocity on the fuselage due to the
presence of the fuselage is of the order of 0.03 in Mach number at a
Mach number of 0.85. The extent of this region of induced velocities
is relatively local, and therefore, for the wing-fuselage configuration,
only the inboard sections of the wing would be significantly affected.

For a Mach number of 0.90 (fig. 3(b)) supercritical conditions exist
over most of the wing semispan and on the fuselage in the region of the
wing-fuselage juncture. It should be noted, however, that based on the
component of velocity normal to wing leading edge the flow over the wing
is still subcritical. The schlieren composite indicates the presence of
weak shock waves in the region above the wing-fuselage juncture. Compar-
ison of this picture with those obtained at other times indicates that
these shocks are extremely transitory in nature. There is no percep-
tible drag rise associated with the formation of these shocks.

At a Mach number of 0.94 (fig. 3(c)), a stronger, extensive shock
stabilizes at the trailing edge of the wing-fuselage juncture as shown
in the schlieren composite. The presence of the shock is indicated by
the dark, shaded region at (a) in figure 3(c). This shock is approxi-
mately normal to the wing surface and, as evidenced by the pressure
distributions, extends laterally normal to the plane of symmetry to
beyond the 60-percent-semispan station. The pressures measured on the
fuselage indicate that the shock emanating from hne wing-fuselage juncture
trailing edge extends with nearly uniform strength around the fuselage.
The nearly normal shock crossing the wing tip ((b) in fig. 3(c)) is
associated with disturbances produced by the tip. The other weak shocks
are transitory. Tuft patterns show no changes in the boundary layer on
the wing or fuselage. The force results indicate that a slight increase
in drag coefficient is associated with the development of the shock on
the wing at this Mach number.







NACA RM L52DOl


The pressure distributions on the fuselage and side-view schlieren
pictures of figure 3(d) indicate that, when the Mach number is increased
to 0.97 the shock originating at the trailing edge of the wing-fuselage
juncture ((a) in fig. 3(d)) becomes relatively weak and slopes rearward.
These same data indicate a strong, nearly normal shock (b) develops
approximately one-half chord length behind the wing-fuselage juncture.
The shading downstream of the strongest indication of this shock (b)
indicates that it curves rearward from the plane of symmetry. Although
the shock (a) is not sufficiently strong to be visible in the lowest
side-view schlieren picture, it probably extends outward to merge with
the shock (b) at approximately (c). The pressure distributions on the
wing indicate that the shock (b) crosses the rearward portion of the
wing and merges with shock (a) in the midsemispan region. The shock
resulting from the merger crosses the outboard region of the wing and
extends nearly normal to the stream well beyond the wing tip, as indi-
cated by the schlieren plan view at (d). The portions of the shock
beyond the wing tips which are nearly normal to the stream are also visi-
ble in the side view at (d). The shading forward of the strongest indi-
cations of the shock (d) in the side view are further indications of the
forward extension of the shock ahead of the normal portions as shown
directly in the plan view. The two dark regions (e) visible in the plan-
view schlieren at a Mach number of 0.97 behind the wing-fuselage-juncture
trailing edge are associated with the nearly normal portions of the com-
bined shock shown at (e) in the lower side-view picture. This relation-
ship is indicated by the shadings ahead of the darkest regions and the
dual nature of the indications. The fact that these indications (e) in
the plan view are not at the same streamwise station as the normal region
at (e) in the side view is due to slight differences in the test Mach
numbers for the two pictures. The double indications are associated
with a slight angle of attack of the model with respect to the stream.
The angle of these dark regions (e) with respect to the stream direc-
tion indicates that at a vertical distance from the combination the shock
is nearly normal to the stream in the spanwise as well as the vertical
direction. This phenomenon results from the rearward slope of the shock
near the plane of symmetry and the forward slope near the tip, as shown
in the side view.

The waves (f) which appear above and below the juncture in the
1
schlieren composite were caused by -inch long, O.02-inch-diameter
wire segments placed normal to the air stream on the upper and lower
surfaces of the fuselage. It has been shown (ref. 4, for example) that
small disturbances in the flow generate waves which cross the flow at
approximately the Mach angle and yet can be detected by schlieren
apparatus. It was hoped, therefore, that the protuberances on the
fuselage would provide an indication of the Mach number distribution in
the region of the wing-fuselage juncture. The protuberances were suf-
ficiently large, however, to produce a strong complex field to points







NACA RM L52D01


at least one diameter from the fuselage surface, and indications of the
Mach number distribution provided by the angles of the waves are inaccu-
rate in that region. Waves emanating from the protuberances become some-
what weaker several diameters from the fuselage surface and do provide a
fairly reliable indication of the Mach number variation at a distance
from the fuselage. The extent of the waves (fig. 3(d)) indicates the
presence of supersonic velocities well above the fuselage surface for
this stream Mach number of 0.97. A measurement of the wave angle at a
point approximately three diameters above the juncture trailing edge
indicated a Mach number of 1.02 in that region.

The pressure data of reference 4 indicate that at this Mach number
of 0.97 the induced Mach number increment on the surface of the fuselage
alone is approximately 0.04. Schlieren photographs obtained at this
condition show that the increased velocity field extends well into the
stream and for the wing-fuselage combination the entire forward portion
of the wing is operating in a Mach number field considerably higher than
the stream value. Also the pressures on the forward portion of the wing
are generally considerably more negative than those for a wing alone.
The pressure distributions and schlieren surveys for the fuselage alone
(ref. 4) indicate that no shock is present on the fuselage alone at this
Mach number; thus, the strong normal shock (b) behind the trailing edge
of the wing-fuselage juncture of the combination must be associated with
the wing. The strength of this shock for the wing in the presence of
fuselage, however, is probably somewhat greater than it would be for a
wing alone.

At a Mach number of 0.99 (fig. 3(e)) the oblique shock originating
at the trailing edge of the wing-fuselage juncture (a) is still rela-
tively weak at the plane of symmetry. At stations farther outboard on
the semispan, the Mach number ahead of the shock and the pressure change
through the shock are greater than those for the inboard region and thus
indicate that this shock is probably somewhat stronger on the outboard
region. When the Mach number is increased to 0.99, the strong, nearly
normal shock present behind the trailing edge of the wing-fuselage junc-
ture at the lower Mach number of 0.97 moves downstream to a position
opposite the tip of the wing ((b) in fig. 3(e)). This shock extends
vertically from the fuselage in a direction nearly normal to the stream.
In the plane of the wing it extends nearly normal to the stream from the
fuselage surface but at a short distance from the fuselage it turns for-
ward, as shown in the plan view at (c). Because of this forward movement,
the shock (b) crosses the rearward region of the outboard portion of the
wing and leaves the tip at (d). The forward movements of the shock (b)
onto the wing tips are also shown in the side view at (c). Just out-
board of the wing tip the weak shock originating at the wing-fuselage
juncture (a) merges with the strong normal shock (d). The combined
shock (e) extends toward the tunnel wall at a moderate angle with respect
to the stream. At short distances above and below the wing plane the







NACA RM L52D01


shock (b) turns rearward slightly in the lateral direction, as shown by
the shading at (f) in the side and plan views.

The combined shock for the model at a Mach number of 0.99 strikes
the tunnel walls at (g) in figure 3(e). The discontinuity in the indi-
cations of this contact in the plan-view pictures is due to a slight
difference in the Mach numbers for the two pictures. The incidence of
the shock on the wall indicates that the flow field of the model has
expanded sufficiently to produce supersonic velocities at the tunnel
wall. The wall pressures indicate a maximum Mach number increment of 0.02
was produced by the model at the wall at near-sonic Mach numbers (ref. 4).

Since a normal shock is present on the fuselage alone at a Mach
number of 0.99 (ref. 4) at the same location as the normal shock (b) on
the fuselage combination, it may be assumed that this shock on the combi-
nation is due in part to the fuselage. The induced velocities ahead of
the shock on the combination, however, are somewhat higher than those
on the fuselage alone because of the expanded field of the outboard
regions of the wing. The shock on the fuselage of the combination,
therefore, is probably stronger than that on the fuselage alone. On
the basis of the pressures measured on the fuselage of the combination
and the shock patterns observed at the lower Mach number of 0.97, it
may be expected that a normal shock similar to that emanating from the
fuselage would be present behind a wing alone and would probably stand
somewhat forward of its location on the combination.

The schlieren composite shows a bow wave, associated with the
deceleration of the local supersonic flow induced by the fuselage,
located about one-half-chord length forward of the wing-fuselage-
juncture leading edge ((h) in fig. 3(e)). Force coefficients indicate
a rather abrupt drag rise for the combination when the Mach number is
increased from 0.94 to 0.99. Since the tufts show only slight changes
in the boundary layer, most of the drag increase is probably due to the
development of strong shocks rather than to separation. For thicker
wings and those with less sweep, the drag increase at high subsonic
Mach numbers is due primarily to separation (ref. 1).

Schlieren photographs obtained at a Mach number of 1.00, but not
presented, indicate that, when the Mach number is increased from 0.99
to 1.00, the strong, nearly normal shock originating from the surface
of the fuselage moves downstream and no longer crosses the tip of the
wing as it does at a Mach number of 0.99. With a further increase in
Mach number to 1.02, this shock continues to move downstream and reaches
the positions shown at (b) in figure 3(f). A comparison of the data
presented for the combination with that for the fuselage alone (ref. 4)
indicates that this shock is farther rearward when the wing is present.
The change in the pressure distribution on the fuselage associated with







NACA RM L52D01


this movement results in a significant increase in the pressure drag
for the fuselage. Thus, although the shock on the fuselage of the combi-
nation appears of the same strength as that on the fuselage alone, the
losses associated with it must be greater. The shock (a) originating
at the trailing edge of the wing-fuselage juncture is still relatively
weak at this Mach number. The most outboard pressure distribution on
the wing indicates an increase in pressure near the trailing edge which
is apparently not associated directly with the shock (a) crossing the
wing semispan. This pressure change is associated with disturbances
originating at the tip. The effect of these disturbances on the field
is shown at (c) in the schlieren side and plan views of figure 3(f).
The various disturbances emanating spanwise from the tip merge at a
short distance from the tip to form a relatively strong shock at (d).
This shock (d) associated with the wing apparently merges with that at
(b) produced primarily by the fuselage several semispans outside the
schlieren view. The bow shock associated with the wing is shown at (e)
in figure 3(f). The second disturbance which appears in the schlieren
composite in the region above the wing-fuselage juncture (f) is the
intersection of the bow shock (e) on the tunnel wall. At the lower Mach
number of 0.99 this shock does not extend to the wall.

With an increase in Mach number to 1.11 (fig. 3(g)), the shock
originating at the trailing edge of the wing-fuselage juncture (a) is
swept nearly to the tip trailing edge. The secondary disturbances
associated with the tip at a Mach number of 1.02 (fig. 3(f)) disappear
at this higher Mach number of 1.11. The bow shock (c) is apparently
attached to the leading edge of the juncture. The shock (b) associated
with the fuselage of the combination moves off the surface of the fuse-
lage as it does for the fuselage alone (ref. 4). The pressure distri-
butions on the rearward end of the fuselage are the same as those on
the fuselage alone. No separation is indicated by the tuft patterns at
this Mach number as at lower Mach numbers. At higher Mach numbers the
shock originating at the wing-fuselage-juncture trailing edge would
move further rearward on the wing and finally to the trailing edge
where it would remain at all higher Mach numbers. The root bow wave
would become more inclined and would reach the leading edge of the wing
at a Mach number of approximately 1.40.


Angle of Attack of 4

At an angle of attack of 4 and a Mach number of 0.85 (fig. 4(a))
the pressure distributions indicate supercritical flow over the forward
portions of the upper surface of the wing. Relatively high Mach numbers
are associated with the negative pressure peaks formed at the upper-
surface leading edge, a value of 1.62 being indicated at the 80-percent-
semispan station. No apparent drag rise is associated with the super-
critical conditions. The schlieren composite indicates weak shocks







NACA RM L52D01


associated with the supercritical velocities in the region above the
wing-fuselage juncture. The weak shock waves shown above the wing tip
are probably caused by disturbances acting parallel to the stream in
the near-sonic velocity field adjacent to the tip. These waves are.
possibly associated with the weak separation near the tip.

At a Mach number of 0.90 (fig. 4(b)), a shock, which appears to
originate at the wing-fuselage-juncture trailing edge and which is
approximately normal to both the wing surface and the plane of symmetry,
is indicated by the schlieren composite ((a) in fig. 4(b)) and the pres-
sure distributions. This shock is similar in appearance to one which
occurred on the wing at an angle of attack of 0 and a Mach number of
0.94. Several disturbances, which seem to emanate from the lower sur-
face of the fuselage in the schlieren composite (b), are associated
with the near-sonic flow indicated on the lower surfaces of the wing
and fuselage by the pressure distributions. A disturbance approxi-
mately normal to the flow located just back of the tip and extending
laterally well beyond the tip is shown in the schlieren plan-form view
at (c). This disturbance may be associated with the deceleration of
a wide accelerated flow field around the wing-fuselage combination. It
is similar in nature to that for a body alone. (See fig. 4(g).)

A definite drag rise occurs when the Mach number is increased to
0.94. The noticeable redirection of the tufts outward on the rear
portions of the upper surface of the wing behind the shock which origi-
nates at the wing-fuselage-juncture trailing edge indicates a thickening
of the boundary layer at this condition which may be associated with
limited separation. It would appear, therefore, that the drag rise
for this condition is due in part to additional boundary-layer losses
as well as to shock losses, unlike the case at an angle of attack of 00
where the drag rise was due almost entirely to shock losses. The
shock (a) extending above the trailing edge of the wing-fuselage junc-
ture in the side-view schlieren composite for a Mach number of 0.94
(fig. h(c)) is considerably stronger than that for an angle of attack
of 0, as might be expected. Conversely, the shock (b) below the junc-
ture is weaker than for 00. The pressures and schlieren surveys indi-
cate that near the tips the shock above the wing is much stronger and
more extensive for 40 than for an angle of attack of 00 (see (c) in
fig. 4(c)). The indications of the shocks above the two tips vary
considerably in intensity and extent. Photographs taken at other
instants for this same condition indicate that the differences between
the strengths of the shock on the two wing tips is due to unsymmetrical
fluctuations of the disturbances since at some instants the shock is
more nearly equal in strength on the two tips. Although it is not per-
ceptible in the composite schlieren photograph, pressure distributions
indicate the presence of an oblique shock associated with the leading-
edge peaks. The pressure distributions indicate that this oblique shock
merges with the juncture trailing-edge shock in the vicinity of the tip.







NACA RM L52DO1


At Mach numbers of 0.97, 0.99, and 1.02 (figs. 4Md) to 4(f)) the
general nature of the shock formation above the wing-body combination
for an angle of attack of 4 is similar to that for 00 at the same Mach
numbers, although the magnitudes and positions of the shocks differ.
The weak oblique shock which apparently emanates from the tip leading
edge at a Mach number of 0.99 ((a) in fig. 4(e)) is probably associated
with the initiation of a tip vortex. The relatively weak oblique
shock (a) ahead of the main fuselage shock (b) at a Mach number of 1.02
(fig. 4(f)) and the complex shock formation (a) above the fuselage at
a Mach number of 1.11 (fig. 4(g)) are associated with the flow over
the fuselage at an angle of attack.

Tuft surveys (fig. 4(f)) show that the shock originating at the
wing-fuselage-juncture trailing edge still leads to a thickening of the
boundary layer at a Mach number of 1.11 although the extent of the
thickened boundary layer is less than at lower Mach numbers when the
juncture shock was farther forward on the wing.


Angle of Attack of 60

At an angle of attack of 60 and a Mach number of 0.80 (fig. 5),
the regions of negative pressure on the upper surface of the sections
near the leading edge become progressively broader from the root to the
tip and suggest the presence of a separation vortex such as that
described in reference 8; the fine nylon tufts are directed outward in
the regions of high negative pressures so that the presence of a leading-
edge separation vortex is again suggested. The heavier yarn tufts,
which extend further into the stream from the model surface, are not
directed outward as much as the fine woven tufts and indicate that the
region of reversed flow of the vortex is quite thin. At the tip
sections the relatively low level of negative pressures on the upper
surface, the relatively poor pressure recovery at the trailing edge,
and the slight outward direction of the tufts are indicative of a
thickened boundary layer over the entire chord.


Angle of Attack of 80

With an increase in angle of attack to 80 at a Mach number of 0.80
(fig. 6(a)), the pressure distributions and tuft surveys indicate a
marked rearward spread and a considerable strengthening of the leading-
edge separation vortex. Complete separation over the wing from a
station just inboard of the 80-percent-semispan station out to the tip
is indicated by pressure surveys.

With an increase in Mach number to 0.85 (fig. 6(b)), tuft patterns
show that outward flow in the boundary layer on the leading edge of most







NACA RM L52D01


of the semispan and over the entire chord on sections inboard of the
50-percent-semispan station has disappeared; this condition indicates
that the leading-edge separation vortex has been eliminated in these
regions. The vortex type of flow has been replaced by an attached
supersonic accelerating flow around the leading edge of the type
described in reference 9.

When the Mach number is increased through 0.90 to 0.99 the extent
of the vortex and separation contract outboard and rearward (figs. 6(c)
and 'Id'i). The tuft patterns and pressure distributions indicate that,
at a Mach number of 0.99, the extent of severe boundary-layer losses
on the wing upper surface is limited to the region back of the adverse
gradients, associated with the shock. At this condition, the flow
phenomena are similar in nature although different in magnitude to the
flow phenomena which existed at a corresponding Mach number and an
angle of attack of 4.

With an increase in Mach number to 1.11 (fig. 6(e)) the shock which
originates at the wing-fuselage-juncture trailing edge moves farther
rearward as it crosses the wing as it did at lower angles of attack.
The pressure distributions indicate that this shock causes a distinct
pressure discontinuity as it crosses the tip section. There is a possi-
bility that the boundary layer is extremely thin over this region of
the tip so that the field pressure disturbance is allowed to extend
nearly undistorted to the model surface.


Angle of Attack of 12

As the angle of attack is increased to 120 at a Mach number of 0.80
(fig. 7(al) the leading-edge separation vortex spreads rapidly rearward
with complete separation over the wing evident from the 30-percent-
semispan station out to the tip.

With increases in Mach number to 0.89 (fig. 7(b)), 0.99 (fig. 7(c)),
and 1.11 (fig. 7(d)), the separated region on the upper surface of the
wing contracts outward and rearward as it did at lower angles of attack.
At a Mach number of 1.11, separation is confined to the region back of
the adverse pressure gradients on the midsemispan and outboard sections.
Because of the reattachment of the boundary layer, the lift carried by
the outboard regions of the wing increases as the Mach number is raised
from 0.89 to 0.99 and, as a'result, the lift of the entire wing increases.
With a further rise in Mach number to 1.11, the lift on outboard sections
continues to increase, while on the inboard sections a decrease in lift
is noted, because of the presence of a supersonic type of flow over
these regions.







NACA RM L52D01


Angle of Attack of 200

At an angle of attack of 200 and at Mach numbers of 0.79 (fig. 81a)'
and 0.89 (fig. 8(b)) severe separation of the flow over the entire upper
surface of the wing is indicated by the tuft patterns. Because of the
complete separation, the negative pressures on each of the sections are
very nearly uniform with the negative pressure level decreasing from
root to tip and indicate more severe separation on the outboard sections.
The greater severity of separation on the outboard sections is also
shown by the tuft patterns. As the Mach number is increased to 0.99
(fig. 8(c)), the pressure distributions and tuft patterns indicate a
reattachment of the flow on the rear portions of the sections near the
root. With a further increase in Mach number to 1.11 (fig. 8(d)), this
region of flow reattachment spreads slightly outward.

At a Mach number of 1.11, the absolute pressures on the inboard
upper surface of the wing approach absolute zero, the limiting pressure
coefficient at this Mach number being about -1.16.


CONCLUDING REMARKS


A study of the pressure distributions, tuft patterns, and schlieren
surveys obtained for a 450 sweptback wing-fuselage combination forms the
basis for the following general remarks.

At angles of attack of 0 and 4 a strong normal shock develops
behind the trailing edge of the wing-fuselage juncture at a Mach number
of 0.97. This shock crosses the wing at Mach numbers of 0.97 and 0.99.
At higher Mach numbers it moves downstream of the wing.

A shock, which originates at the trailing edge of the wing-fuselage
juncture, develops at a Mach number of 0.94. This shock slopes rear-
ward and becomes relatively weak as the Mach number is increased to 1.00.
This shock merges with the strong normal shock behind the wing juncture
and the combined shock extends well beyond the wing-fuselage combination
at near-sonic Mach numbers.

Because of the induced flow over the fuselage, a bow shock forms
somewhat forward of .the wing-root leading edge at a Mach number of 0.99.
This shock moves rearward with increasing Mach number and is at the
root leading edge at a Mach number of 1.11.

'Tuft patterns and pressure distributions indicate no separation
over the wing for all test Mach numbers at an angle of attack of 0.
At an angle of attack of 40 increased boundary-layer losses form between
Mach numbers of 0.89 and 0.94 on the midsemispan and outboard sections







NACA RM L52D01


of the wing back of the shock originating at the wing-fuselage-juncture
trailing edge. With further increases in Mach number, the region of
increased boundary-layer losses contracts rearward.

At an angle of attack of 80 and a Mach number of 0.80, the tuft
patterns indicate the presence of a leading-edge separation vortex and
separation over the midsemispan and outboard sections of the wing. As
the Mach number is increased to transonic values, the flow reattaches
over the forward regions of the wing. At Mach numbers higher than 0.99
separation is confined to the region back of the shock which originates
at the wing-fuselage-juncture trailing edge.

At an angle of attack of 200 and a Mach number of 0.79 the flow
is separated over the entire upper surface of the wing. At Mach numbers
of 0.99 and higher, the tuft patterns indicate a reattachment of the
flow on the rearward regions of the inboard sections.


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






NACA RM L52D01


REFERENCES


1. Whitcomb, Richard T.: An Experimental Study at Moderate and High
Subsonic Speeds of the Flow Over Wings With 300 and 450 of
Sweepback in Conjunction With a Fuselage. NACA RM L50K27, 1951.

2. Edwards, George G., and Boltz, Frederick W.: An Analysis of the
Forces and Pressure Distribution on a Wing With the Leading Edge
Swept Back 37.25. NACA RM A9K01, 1950.

3. DLnforth, Edward C. B., and O'Bryan, Thomas C.: Pressure-Distribution
Measurements Over a 450 Sweptback Wing at Transonic Speeds by the
IIACA Wing-Flow Method. HACA RM L1iD24, 1951.

4. Ritchie, Virgil S., and Pearson, Albin 0.: Calibration of the
Slotted Test Section of the Langley 8-Foot Transonic Tunnel and
Preliminary Experimental Investigation of 3oundary-Reflected
Disturbances. NACA RM L51K14, 1952.

5. Loving, Donald L., and Estabrooks, Bruce B.: Transonic-Wing Investi-
gation in the Langley 8-Foot High-Speed Tunnel at High Subsonic
Mach lumbers and at a Mach Number of 1.2. Analysis of Pressure
Distribution of Wing-Fuselage Configuration Having a Wing of
450 Sweepback, Aspect Ratio 4, Taper Ratio 0.6, and NACA o5A006
Airfoil Section. NACA RM L51F07, 1951.

6. Osborne, Robert S.: A Transonic-Wing Investigation in the Langley
8-Foot High-Speed Tunnel at High Subsonic Mach Numbers and at a
Mach Number of 1.2. Wing-FuEelage Configuration Having a Wing of
450 Sweepback, Aspect Ratio 4, Taper Ratio 0.6, and IACA 65A006
Airfoil Section. NACA RM L50H08, 1950.

7. Loving, Donald L., and Williams, Claude V.: Basic Pressure Measure-
ments on a Fuselage and a 450 Sweptback Wing-Fuselage Combination
at Transonic Speeds in the Slotted Test Section of the Langley
8-Foot High-Speed Tunnel. NACA RM L51F05, 1951.

8. Lange, Roy H., Whittle, Edward F., Jr., and Fink. Marvin P.: Investi-
gation at Large Scale of the Pressure Distribution and Flow
Phenomena Over a Wing With the Leading Edge Swept Back 47.50
Having Circular-Arc Airfoil Sections and Equipped With Drooped-
Hose and Plain Flaps. NACA RM L9G15, 1949.

9. Lindsey, W. F., Dsley, Bernard N., and Humphreys, Milton D.: The
Flow and Force Characteristics of Supersonic Airfoils at High
Subsonic Speeds. HACA TN 1211, 1947.








NACA RM L52DO1


'N.
2~c


S. p


S L-6NACA-3

L L-68VC5


Figure 1.- Wing-fuselage combination mounted on the sting-support system
in the 8-foot slotted test section.


".'.,:y .










NACA RM L52D01


Sting


3.334


F E
450


Figure 2.- Pressure model dimensions and orifice locations.
are in inches.


All dimensions


M.A.C.= 6.125


0.25-chord line


33.333









18 NACA RM L52D01



co
CO





S\ "u I (--





0000
















a
o o 0 5



















I I O in O 5
S, II0 0
II






S'I







r-P









'd
4 \ 4-


















I


r ra)


r IN









NACA RM L52D01 19














co






--
-













rt C

S0



I-I








NACA RM L52D01


Q --
0000 3
0 0 0 "i
oE~a
o o o a~


co0
co


o 0

II









NACA RM L52D01 21
















-- C-


CL








-P
z I















Sr-4
01 )




b.
** "-





w























-ooo S

oocid
I I
0oo0- 3
d d d0 "_


NACA RM L52D01















3\ 0











UIM

/




S*0
o a

I *
ri
o

I 1















Q.. ,O













u rn <








NACA RM L52D01


co

/ -L

N 5 J


r --,-.


JI
v i
*IL.tfT .,
-lagy -s.



?" ^J


I

0

o

0\
0l
























o-oo o.
0008

o o0 "J
000o^^3
O Oa-~l


0. L o
! -


NACA RM L52D01


































-*
0d




O O
o\ -P

U
II
I




*H
cy'


CL -
U 7








NACA RM L52D01 25








co
co




..J
0L
3: W










IId




0
I a





O
4-r


.)

,' 7


U"'


w -























--Oo U

0000

'J 'b "E
':0- I


NACA RM L52D01












07


/ \
o 0












II







o
0* 0

II



a ^
( h
-, a







NACA RM L52D01 27







0


o> > i l-> >>
-- I
Fo


I O- O |







-."
I a

O i"' u
or-I


i,i



> / i


Li
C


!
























--- Or
QQQQ in
0000 '



I
d~ d d1 d
i! a) a>0~ ~


-

0a


NACA RM L52D01

















E-
r r-0














D
,1
SON






















II
!*








NACA RM L52DOI1 29












w
>z
z w















0 |P
.--.T--\ I -. --


r





-J-4
.K- .








30 NACA RM L52D01
















000- QC !5


II '





ap












rj
\'I r-4 *H

-4 o
I











o0 ? o
I

a_


















|~






NACA RM L52D01


0



Lr


.. .m


(. "
S cr

I
e-1


c'

























0-00 S


r'1 Q "E o
U0
I


0 i o



co


NACA RM L52D01




c in

It






cv








rd




II
o














II
II





,0
c*














0)
u



















NACA RM L52DO1


"Y- : .








>
w

cn
Cl)


w

z
-I



K.









"Tr

-&-;:- --


o
rc











o
0-
45







I


+I
*
0)


.1-
(L>


, .
-'u,**

r I 1 I I Ilf l







34 NACA RM L52D01












co -o 8 8

0 D _.

"' "E


















I II







O C'
I I '1





II
000
o 0











i


.Jr




F II


0aC <







NACA RM L52D01


cc
cr
c -
)<


O1





















-ro
C" o !S
0000
I I 3i
E Lo


NACA RM L52D01














ON

t"-






1I





I 7



I
I -t

11


I I
a ^

Ii '> ^


ow U.






NACA RM L52D01


z
a.




LI.

-----------


34

-o


O








NACA RM L52DO1


Q> QU
C\J N C\J LO I S
1- Q o o
OOO d o
I I
'" "o E "- g
I


0 0
C.
"F i


4




I







o a

aI




*-4








NACA RM L52D01


r

w
Ul

w


.J
z
<
D-
a.










r-T
kc


* F,',


i" *.


I..
;IJ
-1r-* n


5

._...i


O ki





















a) 0)
Kror)0
Cd r4) N r) N

ca. 0
S'E "a E "
S0 0 LO D J
8_ 1
^^^a-^


NACA RM L52D01


































I -0 -
i P















I *0



I
o 4


0 -l
-t

a.









NACA RM L52D01


z
-J
Q.





""*H; -^'H --~--- -


p-


w


C
UJ
Q,


I ~--







42 NACA RM L52D01











/ \ \ ro r o rS
rOOQO 0













\I I






LPI








L 00
\ I I I r















\ /
\I
S1 /






SI









I to_L <








NACA RM L52D01


Iry


L
,









44 NACA RM L52D01

















SI
I
00 ro r o z















\ I

I I



H *-4




HO
I I

I i

-I I


\-H
I r F4


f II LL
a -
Q1 m t








NACA RM L52DO1 45









5
>

z


-j
.'






I -. .-






rl
u- -


S-O











LUl
,.,








46 NACA RM L52D01






/1m
00
Z7



o0 roo 0SS



o0

II









\co





*

ON*
\ \I -
\ rd



\ 1 ;o


I I




I I
'z-


0 o IL
0 <








NACA RM L52D01 4"-







"0N
) W 0



C\L to |

Q-0 o

00000













S/\ 1 1 I- :
\I I
I.a




q\ Io





rd


I .

CYO
PA











SI .


II








48 NACA RM L52DO1



II








0 o0 oro 8 (
COOOO e














7^ /I /11 a

E '
\ I









-- \*\i i


9~~~ O hi








NACA RM L52D01 49







I



CD- 0

Ydoo- o
0r-0
b o 0 o I5









S I
I
\ i
Eio\ C1
(5p






*o 4
o
I I


i I







I
I
S'I

So '





II
& \
L ---II- \\ I i u <


om <








50 NACA RM L52D01












*0000









S0I I1
\ *








z ,'i
O \


d
SI\ \
Z __I \ 0


CL
n


uO L.
LDm







NACA RM L52D01 51










I,-o D 8
wo o o- t
'OOi



SI











4 i" \\ I I : a
.I





\ j \ 1 I r- 0-
\ I1
\ 1



I v

JI_ I oI \





i


IP


I I


I




9 I

UM <








NACA RM L52D01


,, in ,n
r -- o 3
0000

00 0 00
o^^o5^^^I
dddI
c~B~dlsI


II











o P
4-P





o 'd
c* 4
0
















I
0
iC





4.,4


tn


4)
PS







Fr,


0 ow L
QOtf. *'









NACA RM L52D01 53













II
Ii









S0I I-
I I i
I I















\0 4-3


II
\ \
\ )r-



I\ \


nI
I..


Om <

















w c
00-00(

", E "E "
00 -0 0 l
d d dd "_I
I ca
o~o~o?


0
7-


NACA RM L52D01










I


i




01/r-1





0 0

I t
I


I I
I'

1 I j
I













V







NACA RM L52D01 55










co o 0
r- r


\I I
\ I I I
'Iooon- L


S"
\\0






/\\


SII I

S\ i I



r-
) \ \ I\



\ I !


om C <








NACA RM L52D01


ooodd

co a E "o 0 0 0 C& =1 -9
i 'bg'11


o r 0o

C


f-








I






6
0.

0






*,
O
Cd







I
p4














e











Cd
0




1
Id
P0
" )









rC4


aum <
im3







NACA RM L52D01 57













\Ij 0
Co rr) tn
S0000 /-
\ .' 'b "E 4l, -
\ \ 0 0 (D-O_ I '-
0\ '





\ \ \\
\\ I


\ 0


\\ \
I




\I o -I
S1 1


II



I















-0 W"




















roOO 0
I I I


NACA RM L52D01




I
I
/











7;-





\ I

a a
a I a o







\ o
ua
Ia
II

1'

II
I I
a I









I a
a af








NACA RM L52DO1 59



I

11



rIn r- N Il
-OOO
'i o



I-



S\
\ \



I\ I I


S\ \ \I I o




-r-t
\ ; U I i
II^^ i:


am <<






NACA RM L52D01


<1 00


,6 b c .,u o 0
a' O to ro c u c -
03 'JUai3jaO3 E)DJG


0 0


-o


E
C
0


-ao V

0
E


CD 0 OD
o o 9 9 ,-
U0 ua!! o u I
wo &JUD1!j18O3 juaWOW-fJU6 pI3d


85 o
04
-s ^


NACA Langley Field, Va.


73 Q ueIOJQ4o
~) 'luapuiaoa un


~rJ


C)
4-3





0


ci-I

H *H
0
o


















E -P
-2

Cd
40
co





r













0
II
-r-t

- Fr


: I




021 i *


-4 CU 02t


00



mtf



0 7



zr






| z "
^ id

z1ze5E~s


0 TO 0 -

CU g3a g
*ao CU -S .
00 Q rV
ca g > 0g
0 4
"0 0
!U o m ...


- a- ia g- L
gCUod -olko


to o o
. og a a a

., e .0 o C
"d 5 0 ..i 0-



- m .. +,0


Si C dv


0u u

o a.

91 d.




0



L i E 2





0 rx.




z z +. o


I~W
a 4a
5 *S2c f. 0
i- ca ra a
0 bD.2 0 M
0 0
3 CU C 0 0a0
S00 a' -c'
- C ^ a
Vt2 a0

cd 0 o 9' 0 0
k'Uto > BDT -


a .a- 7 a 2C


r- a aal ~'
~0CUE c.
cl ra o

Z2 C U
0 CE d 14



0 -a
oU 0 O dl-. 0
rS ^he"" ,

:3 biD B 0a
.2 .2 g4
Pa goa ^ o a a
WoUCU. I 5..~g
9 ao fCU o(
Pm o*-"E^,-3


02 I CO, -
C4i C4 :5 2W F4



Z00 P .0 r E.<


0 1 o -
(Uo bD O2 3
CU 0 -U
g -t c o i .5 w ,,



ma 0 0< w:
Bscllu C 3 z






a 60.2 0

uP4 a a ,c



= r- L .5a s

W1 8a 0 cu 0 C
0 Wo0







r 0
LO -3l C:, gs s


oaa s 2 z < F 0, C -" w
r!g-< S 0 p;g s :
w ,4 a .

z .4.-.20CU2'














tocu~j U ,nU
CUC IV C 0
e d C- C0


.I2~ CU ..0 X
2 i .4 .i t7
1 '^3 >- S1 ''



MOwUE EK U
.04)

-- CU C
; ej ci~ >i B


m 02
uu 3?

E4~m
^ yd 06
.00 .






00
-0g 0
(D CU0
C..
Pbll








IsJt
0j o r
C. IT wo
U)U




0 E40 L) .




=. z
z z 2 E&4 S ^


2 bD 0 0

50 a S 2 ,
- 02 w c S



d w

-0
.4) r- 0w)0
.s- .- 0w g y








S0 ;,. 0 0
, C- o o 4.

i 0 C.. 0 g o cU
4 c o -s oi

4) C. G 0 o C. CU



P402.. 'ii l iC2
1M gO 'S g
ftesa's'sSn


; I t0 =S
pi 2 C- g 4- (i
IC4 u 2 E(
-w -= 0 u c




U cri (6U
u a **^
to co
bo 4) w DOt:3


I: d 0 2 h "
rn u 30 z

Sci c i a


UU
0O S
L) um
20
0 C..
CU 5" .0
wS -0


02EC E ECUa

A4 C:.
w M gd Q6




S03 o5
^^^ o.



0 C.

;o0
CU~

o~~(
s>sls
44z ml


S z-S~ z
am< ug
ZZ&PEN


a t
-0
e 0 b? 00

^: "D o ? E
(D 0 CUC t~U
o a >. .



,U 0 &- .. g o
020 0. .0
O 0r) >.....
2 CU bflC. fC.j..g^. 0

02.0
g '^ji i S u cdi
.0 0
** e q- t kS a) nl
CU 2 k2 C.. 0.
.~ .~ "
*2o gVd ro- E
- 0 Z m 0 0 a)..0
(>L E Ca =g S
'a. 0
. )4) CU Ud





S=u b3 o iCIS
.,~ CU 0 04



C 0 lC.l 0 MlrC
U)0 U~ 0o d.~


g .< 0 CU
U3 i.3.S!


ICU C)


00
C- 1

I z


I3 In
=C)0
C CU
Z
Si r,

c I

^4 C


0uu



* 4 o ,
m p 00






o; a wpE r30 i
0 0.
2ID E- 80i

aoi
UC 0




sog
0U0






". Pki -4 m w 2
C<' g ~4 U




z n4
0 ,o
^*a~ CU
: ^ r... i( g
CU1.U:E
20li
2Z Z4SE-'


C !-


c .f ,., O
5. 5 2 ,d o s
02g 1 M,. o' o



^ 0- 0 4)
U r 3.S o D

0.o C c- 'a
- >,0 04s) CU







0 .0 u .= g.

(D.r.. 00* CU.
0 C. o 0 a
d .0s '- 0 a
20 w0 0 Z
02 0) CUo $
C..0020 CU0o.d
P4 W CUE


ciS

I CU w



via a
t
3l- =
o a
C0
2
u 9--3

CwuE
.. CU U 0
w M S


- CU CU -


00 2



00


20 e.2 0
4i El m .o


SEi. 8-5





S0 r4 K e



Loli

|>84
P4 Z>B Mi ,.
d cc S*'
R L)<'3
o||1
S Cd
z-S z S


c bn 2 0o


Cd14

CU 0


= aU' ) >
0. CU
2 -S 2g .2 c~ CU *
i S 0 6B S ...












S 04) CU
.0 d W ) '






0 > 02
^*a- ".s 's
|s^d9Ca,|"







U C n. U 0..






w CU C 0 0 go CU

00.. CU
r Z -ct
w) U C. CO C r-






P4 to=0 C w
~.- e o h-S c
g.Q g-. s S ( d
Ls c s~a' g S Ba.
m *S oi Soa- g g


CU.0
CU 2 r
-.0 0] CUO


CU .!-o ^_



0
o o
"~ i -.u)
2 E
^1 1

jsl 1^


z


C-SCIC~







. - - - - -- - - -- - -





UNIVLtH i Y ul- I-LUHIU A. .: ;:
I ...... '.. .. .

3 1262 08106 563 2" .y't" "

"




UNIVERSrY OPF LD A
DOCUMENT QEWTIENt
120 MARSTON SacICE .USRAY '
P.O. BOX 117011 :
GAINESVILLE, FL. 3261-1-7011 A,:






-'.-.. .. ". ',
...."..": ? ," *: .' .! :.. .. ; % "..,



="
S ,., .. .., ,':.. ', *, '. .:,T ....








.. : ." .. .
S; .

















IT. .
.: U-. -











i .S '. Oki



.: r 0,. R. I" I
-A .. a.. '., ,"; ii: r : ,
A ;,. ., .,., .' '' E''. i "