Pressure measurements obtained in flight at transonic speeds for a conically cambered delta wing

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TECHNICAL MEMORANDUM jjK

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PRESSURE MEASUREMENTS OBTAINED IN FLIGHT AT TRANSONIC
SPEEDS FOR A CONICALLY CAMBERED DELTA WING

N65 12688

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INASA CR OR TMX OR AO NUMBER)

By Earl R. Keener

High-Speed Flight Station
Edwards, Calif.

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(CATEGORY)

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION
WASHINGTON o October 1959

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NATIONAL AERONAUTICS AND SPACE AHUWISTRATION

TECHNICAL MEMORANDUM X-kQ

PRESSUEE MEASUREMENTS OBIAIKED IN FLIGHT AT TRANSONIC
SPEEDS FOR A CONICALLY CAMBERED DELTA WING*

By Eaxl R. Keener

SUMMARY

\^(o8^

Pressure measurements were made in flight over the conically cam-
bered delta wing of the Convair JF-102A. airplane at Mach numbers up
to 1.19- Maximum angles of attack tested ranged from 2k^ at a Mach num-
ber of 0,70 to 9° at 1.19-

Appreciably large suction pressiores are realized at the leading
edge of the wing similar in magnitude to the high suction pressures
experienced by thin, plane, delta wings. The cambered leading edge is
effective in distributing the low pressures at the leading edge over a
greater frontal area, thus increasing the leading-edge thrust. The
conical distribution of camber results in near-elliptic span-load dis-
tributions at the lower angles of attack; however, a more important
effect of conical camber (together with the wing fences and reflexed
tips incorporated by the JF-102A.) is the delay to higher angles of
attack in the occiorrence of flow separation that normally occurs on a
plane delta wing. A favorable effect on the pressure drag may also be
attributed to the delay in flow separation. Although the outboard wing
fence probably contributes to the delay in flow separation at the tip,
the pressures indicate that the fence induces flow separation inboard
of the fence starting near the leading edge at angles of attack of
about 8^ and extending to the trailing edge as the angle of attack
increases.

A wide variation occvirs in the span-load distributions from a near-
elliptic loading at the lower angles of attack to a near-triangular
loading at the very high angles of attack tested. In general, the dis-
tributions are similar to those of a plane wing, although the delay in
flow separation in the tip region resiilts in slightly larger tip loads.

Title, Unclassified.

• • •

INTRODUCTION

Theoretically, a suction force is predicted along the leading edge
of thin wings at subsonic speeds and also at supersonic speeds if the
leading edge is swept behind the Mach cone. Physical realization of
the suction force results in an appreciable reduction in drag due to
lift- Pressure measurements on thin, plane, delta wings have shown that
a large reduction in pressure, approaching a vacuum, is realized at the
leading edge (ref. l) . However, drag measurements of such wings have
shown that the reduced pressiores do not produce the predicted suction
force because of the small frontal area over which the suction pressures
are distributed (ref. 2).

To distribute the low leading-edge pressures over the maximum possi-
ble frontal area, it was suggested in reference 5 that the leading edge
be cambered. A theoretical study of leading-edge camber for swept and
delta wings (ref. k) showed that, in addition to cajnbering the leading
edge, the span-load distribution must approximate an ellipse to minimize
the induced drag due to lift. A study of surface shapes that result in
an elliptic load distribution led to the development of conical camber.
The amoiint of camber depends on the design Mach number and design lift
coefficient. Wind-tunnel and flight measurements of airplane drag veri-
fied that conical camber results in an appreciable reduction of total
drag at moderate angles of attack (refs. k and 5) •

To study in detail the effects of conical camber on the pressure
distribution and span-load distribution of delta wings, pressure measure-
ments were made in the wind tunnel and in flight. Wind-tunnel pressure
measurements are available at Mach nimbers up to 1.9 in references 6
to 8. The flight measurements are presented herein.

The flight investigation was conducted at the NASA High-Speed Flight
Station at Edwards, Calif., utilizing the 6.3-percent conically cambered
delta wing of the Convair <IF-102A airplane. In addition to conical cam-
ber, the wing also incorporates two fences, a reflexed tip, and an elevon-
control surface. This paper presents an analysis of the flight measure-
ments of wing pressures at Mach numbers up to 1.2. Particular emphasis
is given to the effects of camber on the distribution of the leading-
edge pressures and the effects of the combination of camber, fences,
and reflexed tip on the span- load distributions. In addition, the flow-
separation characteristics, which are not predicted in the theoretical
development of conical camber, are discussed. Comparison is made with
the flight measurements of wing pressures reported in reference 1 for
a plane wing. Tabulated pressure coefficients and integrated aerodynamic
coefficients for all data points are available upon request from the
National Aeronautics and Space Administration.

I« ••• • • • •* •• • ••• • «•«

k^

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1

Cm

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6

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V

%

Cp

^P

Cp, sonic

c

SYl-lBOLS

b/2 wing semi span

l3'/2 wing-panel span, spanwise distance frcm first row of ori-
fices (0.l86b/2) to wing tip

wing-panel bending-moment coefficient about 0b'/2,

/ Cn

Cav "b' b'

wing-panel pitching-racment coefficient about 0.25c,

^ Jo ^'(cav)

'd^
b'

Cn c — ^ >r«^
av D

P - Poo
surface-pressure coefficient, —

Pi " Pu
differential-pressure coefficient, -^^

pressure coefficient for a local Mach number of 1

local wing chord of uncambered section, measiored parallel
to plane of symmetry

c mean aerodynamic chord of wing panel, 2/S / c dy'

-'0

c average chord of wing panel

n wing-section pitching-moment coefficient about 0.25c,

^CP

Cjj^* wing-section pitching-moment coefficient about line per-
pendicular to longitudinal axis of airplane, passing

through 0.25c, c^-h 0.702lfl - ^jc^

r 1

Cj^ wing- section normal- force coefficient.

^0

M free- stream Mach number

p local static pressure

q free-stream dynamic pressure

S/2 area of wing panel (outboard of 0b*/2)

t wing-section maximum thickness

X chordwise distance rearward of leading edge of local chord

X chordwise location of center of pressure of wing section,
lOOfO.25 - ^, percent c

,(0.35 - ^).

chordwise location of center of pressure of wing panel from
leading edge of c, 100(0.25 - -S|, percent c

y* spanwise distance outboard of Ob'/2

y_ spanwise location of center of pressure of wing panel,
lOO(^), percent b72

z local ordinate of wing section, measured normal to chord

line of uncajubered section

a measured airplane angle of attack

6g elevon position

Subscripts:

L left

I lower surface

u upper surface

«> free stream

nESCRIPTION OF MRPIAME AMD WING PAMEL

H A three-view drawing presenting the overall dimensions of the

1 Convair JF-102A airplane is shown in figure 1. Photographs of the air-
1 plane including several views of the wing showing the leading-edge cam-
6 ber, fences, and reflexed tip are presented in figure 2. The physical
characteristics of the airplane and wing panel are given in tahle I.

A drawing of the wing, including cross-sectional views of the
leading edge at each orifice station, is shown in figure 3* The delta
wing has an aspect ratio of 2.08, a taper ratio of 0.023, and zero inci-
dence, dihedral, and twist. The leading edge is swept "back 60.1*^, and
V the trailing edge is swept forward 5^. Wing fences, shown in figures 1
to 5, are located at 0-225bV2 suad O.SOOb'/^. The wing tips are
reflexed 6^ behind the extended hinge line of the eleven (fig. 3). The
wing section is an NACA 000^^—65 airfoil modified by leading-edge camber,
which is distributed conically over the outer 6.5 percent of the local
semispan (fig. 3); the design lift coefficient at a Mach nimiber of 1.0
is 0.166. In providing conical camber the local wing chord was extended
slightly over that for an KACA OOOif-65 airfoil and, as a result, the
wing-section maximum- thickness ratios are slightly less than k percent.
The wing-section coordinates are presented in table II for the static-
pressure-orifice locations.

The geometric characteristics of the elevon, used for longitudinal
and lateral control, are included in table I and figure 3* The fuselage
of the JF-102A is indented according to area-rule considerations for a
Mach number of 1.0.

INSTRUMENTATION AND ACCURACY

Standard NASA instruments were used to record the wing-surface and
differential pressures, indicated free-stream static and dynamic pres-
sures, normal acceleration, angle of attack, angle of sideslip, elevon
^ position, and rolling and pitching angular velocities and accelerations.
The indicated free- stream static and dynamic pressures were obtained
from an airspeed head mounted on a nose boom, and the static-pressxnre
error was determined in flight. Angles of attack and sideslip were

;*

measured by vanes mounted on the nose boom, Elevon position was measured
at the elevon midspan. All instr\aments were correlated by a common timer.

Flush-type static-pressure orifices installed in the left wing were
arranged in seven streamwise rows (fig. 3). The orifices were connected
by tubing through the wing to multicell mechanical manometers in the
instrument compartment. Surface pressures were measured at orifice
rows 1^ 3, 5^ and 7, and differential pressures between the lower and
upper surfaces were measured at orifice rows 2y k, and 6.

Estimated maximum errors of the pertinent recorded quantities and
the resulting coefficients are:

M

±0.02

I

P - Pcx, o^ P^ - Pu^ iVsq ft ±7 I

^e^^ ^^g ±0.2 £

Cp ±0.02

Cn ±0.02

^m ±0.006 ^

% ±0.03

Cm ±0.01

TESTS

The data presented were obtained from pushovers to angles of attack
near zero followed by gradual turns to high angles of attack. The data
cover the Mach number range from O.7 to 1.2 at an altitude of ^4-0^000 feet,

Reynolds number for these tests varied between 23 x 10^ and 58 x 10^
based on the mean aerodynamic chord of the wing.

DATA REDUCTION

Longitudinal control of the JF-102A airplane is obtained by means
of elevons on the wing; therefore, the characteristics of the wing at
zero elevon deflection could not be obtained throughout the lift range.
Consequently, data were selected from the tests at flight conditions
for which the airplane was essentially trimmed at each angle of attack
(near- zero angular velocity and angular acceleration) . Table III pre-
sents the Mach number, angle of attack, and elevon deflection for all
selected data points. For most data points presented, the pressure lag

• ••

resulting from tube length was negligible because the data were obtained
from gradizal maneuvers. Some lag corrections were necessary^ however^
for a few of the low- lift data points obtained from the pushovers to
near- zero angle of attack; these points are indicated in table III. Lag
corrections were detennined by the method in reference 9 for photographic
instruments, and the corrections were checked in flight by comparing pres-
sure measiirements from abrupt and gradual maneuvers. The corrected data
comprised 11 of the total 69 data points and conqpetred favorably to the
zero- lag data.

Automatic digital conrputing equipment was used to obtain pressure
coefficients from the recorded data and to perform the chordwise and
spanwise integrations necessary to obtain the normal-force and pitching-
moment coefficients. Wing-panel coefficients are based on the geometric
properties of the wing outboard of the first row of orifices (0.l68b*/2).

RESULTS

The aerodynamic characteristics of the wing sections are presented
in figures 4 to 6 in the form of curves of normal-force coefficients,
pitching-moment coefficients, and chordwise centers of pressure for nomi-
nal Mach numbers of O.7O, O.9O, 1.02, and I.I9. Figure 7 presents aero-
dynamic characteristics of the wing panel. These figures are presented
as basic data, and only the section normal-force characteristics are
discussed. Chordwise pressure distributions over the upper and lower
surfaces of the wing at four spanwise stations are presented in figure 8
in oblique projection. In addition to Mach niomber and angle of attack,

the pressure coefficients for a local Mach number of l.OfC . ^

V p,sonicy

and the eleven deflection angle are given. The effect of the elevon
deflection may be noted in the pressure distributions by the abrupt
changes in pressure at the elevon junction. The pressixre measurements
are also affected by the outboard fence, by the reflexed tip, and, some-
what, by the elevon-actuator fairing and the inboard fence (fig. 5).

Figure 9 is presented to compare the thicknesswise pressure distri-
bution for an outboatrd section of the cambered wing with that of a simi-
lar wing section of the plane wing of reference 1. Two angles of attack
are shown, 7° and 12°, which represent the moderate angle-of -attack
range . It may be noted that the area under the pressure distribution^
is directly proportional to the wing-section pressiore drag. Consequently,
suction pressures over the forward part of the wing section represent a
negative drag or suction force. In addition to the pressure distribution
over the forward part, figure 9 also includes the pressure distribution
over the rearward part of the upper siorface to show the favorable effect
of camber on the pressure drag in this region. The pressure coefficients

8

are plotted as a function of z/t instead of the usual z/c because
the maximuni thickness of the plane wing is 6.5-P^^cent chord compared
to i^-- percent chord for the cambered wing. Although the distributions
are uncorrected for elevon deflections and some effect of the outboard
fence is present, a qualitative comparison can still be made. For clar-
ity, the leading-edge and trailing-edge locations are noted in the figure.

Figure 10 presents the thicknesswise pressure distributions over
the forward part of the wing section to show in detail the effects of
leading-edge camber on the leading-edge pressures. Distributions are
shown at several angles of attack, and at each angle the pressures at
four spanwise stations are superimposed. In this manner the effects of
increasing amounts of camber can be seen more readily. Since sta-
tion 0.58^b'/2 is located immediately inboard of the outboard fence
(0.600b*/2), the leading-edge pressures include some of the effect of
the fence.

Figure 11 shows a comparison of the wing- section normal-force coeffi-
cient at three comparable stations for the cambered wing and for the plane
wing of reference 1 at M ^ 0.70* The span- load distributions for the
cambered wing are presented in figure 12. Included in the figure are
the locations of the two wing fences, the end of the elevon, and the
reflexed tip. Figure 15 shows a comparison of the span- load distri-
butions with an elliptic-load distribution (with minor axis at 0b'/2,
which is approximately the wing-body junction) . Figure 1^4- compares the
span-load distributions for the conically cambered and plane delta wings.
In this figure the span-load distributions are uncorrected for elevon
deflection for both the plane and the cambered wing, therefore the com-
parison is qualitative only.

DISCUSSION

.Effect of Camber on Leading-Edge Suction

Leading-edge pressures .- The chordwise pressure distributions of
figure 8 show that appreciable suction pressiores are physically realized
at the leading edge of the conically cambered wing of the JF-102A. The
suction-pressure peaks at the leading edge are similar to those for the
plane wing of the XF-92A (ref. l) , For M ^ 1.02 and 1.19, which are
beyond the range of the data for the plane wing, appreciable suction
pressures also occiir, although the minimxjm pressure is generally near
the base of the leading-edge camber rather than at the leading edge as
for the subsonic Mach numbers .

Distribution of leading-edge pressures .- The favorable effect of
leading-edge camber on the distribution of leading-edge pressures may

2G

9

"be seen in figure 9 by conrparing the cambered-wing distributions to the
plane-wing distributions at similar wing stations and at moderate angles
of attack. It is apparent that leading-edge camber distributes the suc-
tion pressures at the leading edge over a greater relative frontal area,
thus increasing the suction force over that of the uncambered section.
For clarity, the increased suction force is represented by the shaded
area between the pressure distributions over the forward part of the
wing sections.

H

1
1
6

From the thicknesswise distributions of figiore 10 for the conically
cambered wing, the favorable effects of camber are first apparent in the
distributions at a ^ 6*^. At a ^ 5° and below, the pressures over
most of the cambered part are greater than ambient, since the angle of
attack of the leading edge is negative to the free stream. At a ^ 6^
the suction pressures are well developed over the leaxiing edge of all
the stations except the root station, 0b'/2, which has no appreciable
camber. The tip station, which has the greatest amo\ant of camber, expe-
riences the most favorable distribution of leading-edge pressures per
unit chord. The tip- station distributions show that it is possible to
distribute the suction pressures over a frontal projection more than
twice the maximum thickness of the airfoil {k percent for the
JF-102A wing) . At the high angles of attack obtained at M «* 0*70
and 0.90 the magnitude of the suction pressures at the outboard wing
sections is reduced as a resxilt of flow separation associated with
wing- section stall.

The favorable effects of leading-edge camber appeeur to continue to
the supersonic speeds tested. At M « 1.02 and I.19 the outboard suction-
pressure coefficients at angles of attack between 6° and 9^ (figs. 10(c)
and (d)) are closer to a vacuxom than for M ^ O.7O and a == 9-0^
(fig. 10(a)). However, it should be mentioned that the angle of attack
for cruise of most aircraft is generally below 3*^ at supersonic Mach
numbers where, for the JF-102A wing, the effects of camber are not bene-
ficial (figs. 10(c) and (d)). It was noted in reference 2 that benefits
of both camber and twist on drag measurements diminish with increasing
Mach nuinber, becoming negligible when the Mach number component normal
to the leading edge exceeded about 0.7* For the JF-102A airplane this
is equivalent to a free-stream Mach number of 1.^4-, slightly higher than
the range of the present investigation -

From the pressure distributions of figure 9 it would appear advan-
tageous to camber the inboard sections of the wing also, even though the
span-load distributions might not be elliptic. However, in this respect
the results of reference 10 are of interest; drag reductions were
obtained for swept wings by increasing the leading-edge radius . Essen-
tially all the drag reduction was obtained by increasing the leading-
edge radius at the outboard stations, which indicates that cambering
the inboard leading edge would, similarly, have little effect on the

10

drag. In fact, figures 8 and 10 show that it may even be detrimental
to camber the inboard sections because leading-edge suction does not
become appreciable at these sections until higher angles of attack are
reached. As a result, the leading-edge pressures at the inboard sections
probably would be greater than ambient pressure at the moderate angles
of attack at which leading-edge suction is desired; thus the drag would
be increased.

Flow-Separation Characteristics

Leading- edge- separation vortex .- For thin wings with low aspect
ratios and highly swept leading edges, the flow over the leading edge H

characteristically rolls up into a vortex, referred to as the leading- ]_

edge- separation vortex. The vortex originates near the wing tip and X

moves inboard along the leading edge with increasing angle of attack, 6

trailing off the wing near the tip and predominating over the tip vor-
tex. Flight data (ref . l) showed the existence of the vortex at Mach
numbers up to 0.93^ and wind-tunnel results from references 6 and 11
show a vortex at supersonic speeds for both plane and cambered wings
with subsonic leading edges. u

In figure 8 the trough in the pressure distributions near the
leading edge indicates the presence of a leading- edge- separation vor-
tex for the cambered wing. The effects, however, are not as prominent
as for the plane wing. The presence of the vortex is expected, even
with leading-edge camber and fences, since experiments have shown that
the vortex has considerable strength. However, the camber and fences
apparently resist the formation of the vortex and delay its effect to
higher angles of attack.

Wing-section stall .- The individual contributions of conical camber,
fences, and reflexed tip cannot be determined from the data herein. The
combined effect on the wing-section-stall characteristics is best eval-
uated by examining the section normal-force-coefficient curves of fig-
ure k and the comparison of the section normal-force coefficients of the
cambered wing euid the plane wing in figure 11 . In figure k the high
lifting efficiency of the two outermost wing sections is readily apparent.
The comparatively high normal-force coefficients and angles of attack
attained before the occurrence of section stall on the cambered wing
are in contrast to the early loss in lift at the tip sections reported
for the plane wing in reference 1 and shown in figure 11. The stalling
of the tip sections of the plane wing at low angles of attack is a con-
sequence of the formation of the leadlng-edge-separatlon vortex. There-
fore, an important effect of the combination of camber, fences, and
reflexed tip is to delay early flow separation at the outboard sections
to higher angles of attack (above 8^, fig. k) . ,

• ••• • • •

• •• • • Xj

11

The delay in flow separation at the outboard wing sections would
be expected to have a favorable effect on the pressure drag. It is diffi-
cult to show this quantitatively by using the plane- and canibered-wing
data because of the wing fences and the differences in the elevon posi-
tion; however^ figure 9 shows an indication of this source of drag reduc-
tion at a =^ 12*-*. The thicknesswise pressure distribution over the rear
part of the plane and cambered wing sections shows a reduction in pres-
sure drag for the cambered section as a result of the delay in flow
separation.

H Effect of outboard fence .- A noticeable effect of the outboard

1 fence (0.600b'/2) on the wing sections just inboard of the fence may be
1 noted in figure k. At orifice station 0.58^b'/2 a large reduction in
6 normal- force- curve slope occurs at an ajigle of attack of about 8^ through-
out the Mach number range tested. The pressure distributions in figure 8
show a reduction in loading at the leading edge between angles of attack
of 8*^ and 10^. As the angle of attack increases above 12^, it becomes
apparent that flow separation is occurring inboard of the fence extending
> eventually over the full chord, as indicated by the lack of pressiore
recovery at the trailing edge. The flow separation inboard of a wing
fence has been shown to be a normal occiorrence for fences on swept and
♦^ delta wings (ref. 12); however, this undesirable effect is usually com-
pensated by the contribution to the delay in flow separation outboard
of the fences.

Span-Load Distributions

The span- load distributions in figure 12 show a wide variation in
loading from a near-elliptic loading at the lower angles of attack to a
near- triangular loading at the very high angles of attack. At moderate
angles of attack, beginning at a « 9^ in figure 12, the effect of the
wing fences on the loading is apparent. The outboard fence resiilts in
a significant reduction in loading on the inboard side as a result of
the previously mentioned local- flow separation. This is contrary to
the theoretical effect which, in the absence of flow separation, should
be to increase the loading inboard and decrease the loading outboard
(ref. 15). Some effect of the inboard fence is also noticeable, with
the fence increasing the loading inboard and decreasing the loading out-
board, as predicted by theory. At a ^ 2k^ (M «= 0.70) the fence effects
are no longer noticeable as a result of extensive flow separation out-
board of about 0.5b'/2 (see also figs. 4(a) and 8(a)); consequently, the
span loading is nearly triangular.

Since the basic purpose in distributing the leading-edge camber in
a conical manner along the span is to obtain an elliptic span-load dis-
tribution at moderate angles of attack, the distributions from figure 12
• at a =« 7° are compared to an elliptic distribution in figi:ire 13 . From

12

tt

this figure it may "be seen that the loading falls generally along the
line for an ellipse. At Mach niambers "below 1.19 "the reflexed tip
decreases the tip loading below the elliptic loading. As a resiolt, the
spanwise position of the center of pressure in figure 7(e) is about
2 percent farther inboard for subsonic Mach numbers than for M ^ 1.19*

In general, the distributions in figure 12 at M ^ O.7O do not
differ greatly from those for a plane wing at low and high angles of
attack, as can be seen in figure ik . At a ^ 9^ and 13^ the previously
mentioned delay in flow separation for the cambered wing results in
slightly larger loads in the tip region than for the plane wing. H

1

In reference Ik the span- load distributions of the JF-102A were 1

compared to those predicted by linear theory for a flat-plate wing of 6
the same plan form. The distributions compared well at the lower angles
of attack, primarily because theory predicts a near-elliptic loading for
plane triangular wings. At high ^.ngles of attack the comparison breaks
down because of flow separation, the effect of which can be seen at
a « 24° in figure Ik herein. Using the theory in reference I5, which
accounts for the vortex at the leading edge, and correcting for elevon
deflection by the method of reference 16, the results are still unsatis-
factory. However, this theory is for wings of very low aspect ratio and,
apparently, shoixld not be applied for any other case. For moderate
aspect ratios there does not appear to be a method available that will
predict the effects of the leading-edge-separation vortex.

CONCLUDING REMARKS

Pressure measurements were made in flight over the conically cam-
bered delta wing of the Convair JF-102A airplane at Mach numbers up
to 1.19. Maximum angles of attack tested ranged from 2k^ at a Mach
niomber of X).70 to 9° at 1,19.

Appreciably large suction pressures are realized at the leading
edge of the conically cambered delta wing similar in magnitude to the
high suction press\ires experienced by thin, plane, delta wings. The
cambered leading edge is effective in distributing the low pressirres
at the leading edge over a greater frontal area, thus increasing the
leading-edge thrust. The conical distribution of camber results in
near-elliptic span-load distributions at the lower angles of attack;
however^ a more important effect of conical camber (together with the
wing fences and reflexed tips incorporated by the JF-102A) is the delay
to higher angles of attack in the occurrence of flow separation that
normally occurs on a plane delta wing. A favorable effect on the pres-
sure drag may also be attributed to the delay in flow separation.
Although the outboard wing fence probably contributes to the delay in

• ••

• • ^. .«•«'

• ••

13

flow separation at the tip^ the pressures indicate that the fence induces
flow separation inboard of the fence starting near the lesuiing edge at
angles of attack of about 8*^ and extending to the trailing edge as the
angle of attack increases.

A wide variation occurs in the spgtn-load distributions from a near-
elliptic loading at the lower angles of attack to a near- triangular
loading at the very high angles of attack tested* In general, the dis-
tributions are similar to those of a plane wing, although the delay in
flow separation in the tip region results in slightly larger tip loads.

a
1

High-Speed Flight Station,

National Aeronautics and Space Administration,
Edwards, Calif., May 5, 1959-

Ik

• • •• •

• • •

• • ••

• •t •

• * • ••

1. Keener, Earl R., and Jordan, Gareth H.: Wing Pressure Distributions

Over the Lift Range of the Convair XF-92/V Delta-Wing Airplane at
Subsonic and Transonic Speeds. MCA RM H55GOT, 1955-

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Ratio Wings at Subsonic and Supersonic Speeds. MCA RM A53A50,
1955*

3. Jones, Robert T.: Estimated Lift-Drag Ratios at Supersonic Speed.

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k-, Boyd, John W., Migotsky, Eugene, and Wetzel, Benton E,
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: A Study of
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1956.

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Supersonic Speeds on a Thin Conical Cambered Low -Aspect -Ratio Delta
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1957-

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8. Phelps, E. Ray: Pressure Distributions at Mach Numbers of 1.6 and I.9

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out Pylon-Mounted Engine Nacelles. NACA RM A56BO5, 1956.

9. Huston, Wilber B.: Accioracy of Airspeed Measiirements and Flight

Calibration Procedures. NACA Rep.. 919^ 19^8.

10. Evans, William T.: Leading-Edge Contours for Thin Swept Wings:
Analysis of Low- and High-Speed Data. NACA RM A57B11, 1957-

An

11. Spahr, J. Richard, and Dickey, Robert R.: Wind-Tunnel Investigation
of the Vortex Wake and Downwash Field Behind Triangular Wings and
Wing-Body Combinations at Supersonic Speeds. NACA RM A53D10, 1953-

15

12, Haines, A. B., and Rhodes, C. W.: Tests in the R.A.E. 10 Ft. x 7 Ft.
High Speed Tunnel on a 7-5^ Thick, 50° Swept Wing Fitted With Stall
Fences and a Leading -Edge Chord -Extension. Tech. Note No. Aero 2521,
British R.A.E. , Sept. 195^.

15. Weber, J.: Theoretical Load Distribution on a Wing With Vertical
Plates. R. & M. No. 296O, British A.R.C, 1956. (Supersedes
R.A.E. Aero 2500, Mar. 195^.)

Ik. Malvestuto, Frank S., Cooney, Thomas V., and Keener, Earl R.: Flight
Measurements and Calcxilations of Wing Loads and Load Distributions
at Subsonic, Transonic, and Supersonic Speeds. NACA RM H57E01,
1957.

15. Brown, Clinton E., and Michael, William H., Jr.: On Slender Delta

Wings With Leading-Edge Separation. NACA TN 3^50, 1955-

16. Diederich, Franklin W., and Zlotnick, Martin: Calculated Spanwise

Lift Distributions and Aerodynamic Influence Coefficients for
Swept Wings in Subsonic Flow. NACA TN 5^76, 1955-

• • ••
• • •

Id * • • • •

• •••

• • •

• ••

• •

• • •••

TABLE I
TABLE OF PHYSICAL CHARACTERISTICS

Wing:

Total area, sq. ft 695. 05

Span (actual), ft 38-17

Airfoil section NACA OOOU-65

(Modified)

Conical camber, percent local semispan . 6.3

Mean aerodynajnic chord, ft 23.76

Aspect ratio 2,08

Root chord, ft 35-63

Tip chord, ft O.8I

Taper ratio O.O23

Sveep at leading edge, deg 60.I

Sweep at trailing edge, deg -5

Incidence, deg

Dihedral (uncambered chord line), deg

Geometric twist, deg -

Tip reflex, deg -6

Wing panel (outboard of wing station 3-5^2 ft) -

Area (one panel) sq. ft 232. 50

Span (one panel), ft 15-52

Mean aerodynamic chord (wing station 8,210 ft), ft . 20*6^

Elevons :

Area (total, rearward of hinge line), sq_ ft 67-2

Span (one elevon), ft 12.89

Vertical tail :

Airfoil section NACA 000^^-65

(Modified)

Area (above waterline 33), sq ft 68.3

Aspect ratio • - • I-l

Sweepback of leading edge, deg 6O.O

Sweepback of trailing edge, deg -5

Fuselage :

Length, ft 63.3

Maximum diameter, ft u.5

Eq.uivalent-body fineness ratio , 9-1

Power plant ;

Installed static thrust at sea level, lb 8,800

Installed static thrust at sea level (with afterburner), lb 13,200

Test center -of -gravity location, percent -mean aerodynamic chord . . . . 28 to 29

3G

• ••

17

TABI£ IT
LOCATIOHS OF STATIC-PRESSURE CRIFICES
[stations and ordinates, percent chord]

VO

I

t

72

0,168

b»/2

0.320

b'/2

0.441

V/2

0.584

b'/2

0.713

b./2

0-851

v/2

station

Ordinate

Station

Ordinate

Station

Ordinate

Station

Ordinate

Station

Ordinate

Station Ordinate Station

Ordinate

Upper surface _}

.

-0.314

2.5

-595

-0.93^

2.5

.130

-2.081

2.5

-1.688

-6.919

0.5

.217

5-0

.848

0.5

-.307

5.0

.616

0.5

-1.367

5-0

-.849

2,5

-5.016

1.0

.380

9-0

1.090

1.0

-.054

N5

.866

1.0

-1,096.

10.0

-35^

5.0

-4-006

2.0

.583

10.0

1.136

2.0

■SOO

9-3

1.011

2.0

-.674

15-0

• 951

10.0

-2.185

3.0

■717

13.^

1.275

3.0

.556

13-2

1.166

3.0

-.333

23.7

I-38O

15.0

-761

k.o

.809

15.0

1-333

4.0

.664

15-0

1.271

4.0

-046

30.0

1.539

20.0

.298

5.0

•892

20.0

1,500

6.0

-655

20.0

1.442

6.0

.383

40.0

1.716

29.4

1.291

7-5

1.05^4

25.0

1.623

7.9

.976

24.8

1.567

8.0

.674

50.0

1.800

42.9

1.589

11.1

1.220

30.0

1.722

10.8

1.L21

29-8

1-677

10.0

,872

60.0

i.809

52.9

1.688

ia.5

1.277

34.8

1-798

12.5

1.204

35.0

1-757

12.0

1.017

71.8

1.688

61.3

1.887

15-0

1.369

40.6

1.866

15.3

1.320

40.0

1.622

17.1

1.262

f9.o

1-446

JO.O

2.715

20.1

1.526

45-2

1.903

20.0

1.474

45.0

1.867

20.0

1-35^

80.0

1.399

80.0

3.526

25..0

1.640

50.1

1.924

25-0

1.598

50.0

1.892

25.0

1,499

85.0

1.100

91 -0

3.989

30.0

1.737

55-0

1.927

30.0

1-702

55.0

1-897

30.0

1.605

90.0

-755

95.0

4.138

35-0

1.815

60.0

i.914

35-0

1-777

60.0

1.887

35.0

1.691

95.6

.364

i^o.o

1.875

65-0

1.873

4o.O

1.843

65.0

1.847

40.0

1-764

k3^o

1.917

70.1

1.780

45.0

1.889

70.7

1.752

45.0

1.810

50.3

1-935

75.0

1.634

50.0

1.910

74.9

1.62?

50.0

1.836

55-7

1.935

80.0

1.398

55.0

1.914

83.0

1.211

55-0

1.843

60.8

1.917

86.8

.947

60.0

1.901

87.5

.906

58.3

1.843

65-0

1.880

89.1

-783

65.0

1.864

90.0

.731

66,1

1.797

70.0

1.789

90.0

.721

70.0

1.777

92.5

-556

70.0

1-731

75-0

1.635

92.0

.581

75.2

1.615

94.5

-^15

75.0

1-599

81. i4

1.312

9^.0

.441

79.9

1-399

97.6

.195

85.2

I-063

85.0

1,069

96-0

.301

84.5

1.121

87.0

-938

88. U

.829

9B.U

.133

87-9

.876

90.0

.733

90.1

.709

BB.^

-830

94.8

.403

91-0

.646

90,0

.722

97-0

.244

93-0

.506

93.0

.516

95-0

.366

96.0

-303

97.0

,226

98.0

.162

^8.6

.112

Lover svx

face

0.5

-0.403

2.5

-.677

0.5

-1.017

2.5

-1.286

0.5

-2.219

2.5

-3-329

2.5

-6.820

1.0

-.440

5.0

-.762

1.0

-1.009

5-0

-1.076

1.0

-2.219

5.0

-2.816

5.0

-6.191

2-0

-.500

9.0

-1.039

2.0

-.967

7-5

-.991

2.0

-2.100

10.0

-1-977

10.0

-4.767

5-0

-.568

10.0

-1.104

3.0

-.897

9.3

-1.021

3.0

-1.942

15.0

-1.473

15.0

-3.476

I4.0

-.646

13.^

-1.268

4.0

-.872

13.2

-1,156

4.0

-1.790

23.7

-1.380

20-0

-2.400

5.0

-731^

15.0

-1.333

6.0

-.868

15-0

-1.271

6.0

-1.506

30.0

-1-539

29.4

-1.440

/-5

-.96y

^0.0

-i.ijoo

7.9

-.922

20.0

-1.442

8.0

-i.jOo

40.0

-1.716

42.9

-1.589

11.1

-1.197

25.0

-1.623

10.8

-1.100

24.8

-1.567

10.0

-1.189

50.0

-1.800

52.9

-1.688

11.9

-1.260

30.0

-1.722

12.5

-1.191

29.8

-1.677

11.7

-1.143

60.0

-1.609

61. 3

-1.556

15.0

-1.366

34.8

-i.798

15-3

.1.312

35.0

-1.757

18.3

-1.328

71.8

-1.688

70.0

-.613

20.1

-1.526

40.6

-1.866

20.0

-1.474

40.0

-1.822

20.7

-1,420

79-0

-1.446

60.0

,712

2^.5

-1.640

45,2

-1.903

25-0

-1.596

45.0

-1.867

25.0

-1.539

80.0

-1.399

91-0

2,500

30*0

-1.737

50.1

-1.924

30.0

-1.702

50.0

-1.892

30.0

-1.645

85.0

-1-100

94.6

3.178

35,0

-1.815

55.0

-1.927

35-0

-1.777

55-0

-1.897

35.0

*-2.642

90.0

--755

39.6

-1-875

60.0

-1.914

4o.o

-1.843

60.0

-1.887

40.0

*-3.230

95.6

-.364

i+5-0

-1.917

65.0

-1.873

45.0

-1.889

65.0

-1,847

45.0

*-3.673

50.3

-1-935

70.1

-1.780

50.0

-1.910

T0.7

-1.752

50.0

*-4.075

55.7

-1.935

75.0

-1.634

55.0

-1.914

74.9

-1.627

55.0

*-4.426

60,8

-1.917

80.0

-1-398

60.0

-1.901

83.0

-1.211

60.0

*-4.723

65.0

-1.880

86.6

-.950

65.0

-1.864

8f.5

-.906

65.0

*-5.ooo

70.0

-1,789

89.1

-783

70.0

-1.777

90.0

-.731

70.0

*-5,284

75-0

-1.635

90.0

-.721

74.9

-1.615

92.5

-.556

75.0

*-5.542

81. U

' -1-312

92.0

-.581

80.0

-1.399

94,5

-.415

85,0

*- 5.040

85.0

-1.069

94.0

-.441

85.1

-1.063

97-6

-.195

87.0

♦-4.346

88.4

-.828

96.0

-.301

87.9

-.876

90.0

*-3-237
*-1.632

90.1

--709

9d-k

-.133

88.5

-.830

9^.9

91-0

-.652

90.0

-.722

97.0

*-.9l2

93-0

-.486

93.0

-.515

95-0

-.366

96.0

-.303

97.0

-.226

98.0

-.162

98.6

-.112

♦Orifices located on exirface of elevon-actuator fairing.

18

S I .

• • ••

• • • • •

• • • ••• ••

' I • ! .! ! !

• • • • •

P >

% I

TABLE III
FLIGHT CONDITIONS AT WHICH PRESSURE DISTRIBUTIONS WERE OBTAINED

M

a, deg

6e. deg

M

a, deg

5e' d^S

M ^ 0.55

M ^ 0.95

0.53

7.0

1.6 up

0.95

-0.5

1.5 dovn

.51

8.2

2.0

.95

6.1

.4 up

•59

26.1

12.2

.95

7.U

2.0

.^6

27.4

fll-O

.95

9.5
12.6

4.6

•95

10.2

M % 0.70

M Z 0.98

>^0.7O

3.0
3.8

O.U up

.5

X-

70

0.98

3-8

. 5 down

X-

TO

k,l

.6

.98

6.0

4.0 up

70

7.3

.9

.98

8.7

6.9

70

9.0

1.4

.99

10.1

8.7

K-

70
69

9.9

2.1
2.4

.97

10.8

t8.7

11.7

66
73

13.1

3.9

M ^ 1.02

15./
18.5

5.3
7.6

*1.02
*1.02

1.0
2.0

1 . h dovn
1-3

• 9 up

i*.5

69
68

22.2
23.7

10.9
t9.6

1.02
1.02
1.02

3.9
6.4
7.6

5.9

M ^ 0.80

1.02

8.8

8.0

1.02

9.U

Q.h

0.82

5.2

1.5 up

.82
.82

6,h

8.9

10.5

2.1
2.3
3.1

M Z 1.10

.80

1.11

1.2

2.2 down

.80

12.0

3.9

1.11

1.7

■ 9

.81

13.0

^.5

1.12

2.8

.6 up

.80

1^.5

5.7

1.12

^.3

3.1

.80

17.8

t4.4

1.10

5.0

1+.8

.80

20.6

8.6

1.12

6,0

6.8

1.11
1.08
1.08

6.8
7.^
8.7

7.6

7.9

9-9

10.3

M ^ 0.90

0.90
^ .90

-1.0

.1

1.7 dovn
1.1

1.07

9.2

<■ .90
^ .90

1.1
2.0

1.0

1.0

M Z 1,19

^ .90

2.9

.7

1.19

0.7

1.1 dovn

^ .90

3.9

.6 up

1.19

3.1

2.5 up

.89

5.1

.5

1.19

7.0

9-7

.92

6.9

1.4

1.18

8.6

11.1

.92

9.0

tl.5

.88

9.7

3.3

.89

11.3

t2.3

.90

12.7

3.9

.90

17.9

10.3

I

^ Corrected for pressure lag.
"t" 6e < 5e ^o^ trim.

>• #•• •

»• ••• ••

!• •• • ••4

19

H
I

^— t

— 458-
(actual )

Figure 1.- Three-view iraving of the airplane. All dimensions

in inches.

20

• • • •

• • •

• 9
• •• •

• • ••• ••

• • • • • •

• • •• • •

• • • • •
••• •• ••• ••

41374

^

E-2554

(a) Coirrplete airplane; side and overhead views. E-2551
Figure 2.- Photographs of the JF-102A airplane.

^mmmmmm^-

• ••• •

• • • «

• •• •

«• •* • •••

21

E-42i+6

E-k2kh

(b) Close-up views of wing showing leading-edge camber, fences,
reflexed tip, and elevon actuator fairing.

Figure 2.- Concluded,

22

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58

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