Langley full-scale tunnel investigation of the factors affecting the directional stability and trim characteristics of a...

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Material Information

Title:
Langley full-scale tunnel investigation of the factors affecting the directional stability and trim characteristics of a fighter-type airplane
Alternate Title:
NACA wartime reports
Physical Description:
30, 64 p. : ill. ; 28 cm.
Language:
English
Creator:
Sweberg, Harold H
Guryansky, Eugene R
Lange, Roy H
Langley Aeronautical Laboratory
United States -- National Advisory Committee for Aeronautics
Publisher:
Langley Memorial Aeronautical Laboratory
Place of Publication:
Langley Field, VA
Publication Date:

Subjects

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

Notes

Summary:
Summary: Tests were made in the Langley full-scale tunnel of the Grumman XF6F-4 airplane in order to investigate the factors that affect the directional stability and trim characteristics of a typical fighter-type airplane. Eight representative flight conditions were investigated in detail. The separate contributions of the wing-fuselage combination, the vertical tail, and the propeller to the directional stability of the airplane in each condition were determined. Extensive air-flow surveys of sidewash angle and dynamic-pressure ratio along a line coincident with the rudder hinge line were made for each condition investigated to aid in evaluating the slipstream effects. The data obtained from the air-flow surveys were also used to investigate methods for calculating the contribution of the vertical tail to the airplane directional stability. The results of the tests showed that, for the conditions investigated, the directional stability of the airplane was smallest for the gliding condition with flaps retracted and was greatest for the wave-off condition with flaps deflected 50°. The variation of sidewash angle at the vertical tail with angle of yaw was destabilizing for all conditions investigated. Propeller operation increased the magnitude of the destabilizing sidewash but, at small angles of yaw, also increased the dynamic pressure at the vertical tail sufficiently to make the combined effect stabilizing.
Bibliography:
Includes bibliographic references (p. 23).
Statement of Responsibility:
by Harold H. Sweberg, Eugene R. Guryansky, and Roy H. Lange.
General Note:
"Report no. L-109."
General Note:
"Originally issued November 1945 as Advance Restricted Report L5H09."
General Note:
"Report date November 1945."
General Note:
"NACA WARTIME REPORTS are reprints of papers originally issued to provide rapid distribution of advance research results to an authorized group requiring them for the war effort. They were previously held under a security status but are now unclassified. Some of these reports were not technically edited. All have been reproduced without change in order to expedite general distribution."

Record Information

Source Institution:
University of Florida
Rights Management:
All applicable rights reserved by the source institution and holding location.
Resource Identifier:
aleph - 003614294
oclc - 71252173
sobekcm - AA00006279_00001
System ID:
AA00006279:00001

Full Text

~jjic u-(C:9


ARR No. L5H09


NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS





TWARTIIME REPORT
ORIGINALLY ISSUED
November 1945 as
Advance Restricted Report L5H09

LANLEY FULL-SCALE TUNNEL INVESTIGATION OF THE FACTORS
AFFECTING THE DIRECTIONAL STABILITY AND TRIM
CHARACTERISTICS OF A FIGHTER-TYPE AIRPLAm E
By Harold H. Sweberg, Eugene R. Guryansky,
and Roy H. Lange

Langley Memorial Aeronautical Laboratory
Langley Field, Va.






NAC "



WASHINGTON

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


L 109


DOCUMENTS DEPARTMENT


C
s



































Digitized by the Internet Archive
in 2011 with funding from
University of Florida, George A. Smathers Libraries with support from LYRASIS and the Sloan Foundation


http://www.archive.org/details/langleyfullscale00








UACA ARP' No. L5-O9 RESTRICTED

NATIOI:AL ADVISORY COr.,ITTEE FOR AERONAUTICS


ADVATNC3E TESTRICTID REPORT


LA GLEY FULL-SCALE VT.I"EL INVESTIGATION OF THE FACTORS

AFFECTION THE DIE'ETIONAL STABILITY AND TRTUT

CHARACTERISTICS OF A FIGHTER-TYPE AIRPLANE

By Harold H. Sweberg, Eigene R. Guryansky
end R:,y H. Lange

SU .' A R Y

Tests were made in the Langley full-scale tunnel of
the Jrumrr-iLn XF-l' sirplanc in order to investigate the
factors tK,.t affect the directional stability and trim
characteristics of a: typical fighter-type airplane. Eight
represent native flight conditions were investigated in detail.
The separate contributions of the wing-fuselage combination,
the vertical tail, ?.nd tie propeller to the directional
stability of the sirrnlane in each condition were determined.
Extensive air-flow surveys of sidewash angle and dynamic-
pressure r-'tio 5o1ng line coincident with the rudder
hinge line were imade for each condition investigated to
aid t~. evaluatin- the slipstream effects. The data obtained
from the ?.ir-flo'.. surveys 'ere also used to investigate
methods for calculat-in the contribution of the vertical
tail to the airplane directional stability.

The results of the tests showed that, for the condi-
tions I-.nvtigdsted, the directional stability of the air-
plane was smallest for the gliding condition with flaps
retracted arid was greatest for the wave-off condition with
flaps deflected 530. The variation of sidewash angle at
the vertical tail with angle of yaw was destabilizing for
all conditions investigated. Propeller operation increased
the magnitude of the destabilizing sidewash but, at small
angles of yaw, also increased the dynamic pressure at the
vertical tail sufficiently to make the combined effect
stabilizing. The lateral displacement of the slip-
stream with respect to the vertical tail at angles of yaw
larger than annroxirmately 100 caused a reduction in the
contribution of the vertical tail to the airplane direc-
tional stability at positive angles of yaw-and an increase


RESTRICTED










2 NACA ARR No. L5H09


at restive angles of yaw. Flao deflection tended to
increase the directional stability of the airplane
regardless of the condition of propeller operation.

The rudder deflection required for directional trim
was greatest for the wave-off condition with the flaps
deflected 50'. The large changes in the directional trim
of the airplane resulting from propeller operation are
primarily due to the effects of the slipstream on the
wing-fuselage combination and on the vertical tail and
are only secondarily due to the direct effects of the
propeller forces.


INTR DUCTI 0


The importance of the effects of propeller operation
on the directional stability and trim characteristics of
an airplane is well known. Past experience has lhown
that the directional trim is usually critical for a take-
off or low-speed climb condition in which high propeller-
thrust and torque coefficients produce large increments
of ys-.'ir.i-moment coefficient. For such conditions, a
pilot may often find that, because of the large trim
char-es involved, he has insufficient rudder control and
is unable to maintain the desired heading. The directional
stability is usually lowest for a condition of high ani.le
of attack and low power, during which the contribution
of the vertical tail to directional stability is lowest
because of the low slipstream velocity and the relatively
large loss in dynamic pressure due to the fuselage and
canorp wakes.

Analyses have been made in the cast of w,"ind-tunnel
data on directional stability and control (references 1
and 2) but these analyses were based mainly on the
results of scattered tests of a large number of airplanes
anrd. airplane r-odals and did not include any systematic
test results showing the effects of propeller operation
on the directional stability and control characteristics
of a single design. In particular, only meager data were
available to show the effects of proFpller operation on
the air flow in the region of the vertical tail. In
order to obtain some systematic wind-tunnel-test data
relative to these effects, an investigation was conducted
in the Lar: ley full-scale tunnel on the Grumman XFbF-1
airplane. The investigation included measurements of the










NACA AFPR No. L5HCQ


directional stability and control characteristics of the
airplane for a wide range of flight conditions. For each
flight condition investigated, tests were made of the
complete airplane, of the airplane without propeller, of
the airplane without vertical tail, and of the airplane
without both propeller and vertical tail. The separate
contributions of the propeller, the vertical tail, and
the wi.ng-fuselage combination to the airplane directional
stability and trim could thus be evaluated. In addition
to these force tests, measurements were made of the
dynamic pressure and the angularity of the air flow at
the vertical tail. Particular attention was given to
these air-flow measurements inasmuch as the available
data on this subject are very limited.


SY MBO LS


CL lift coefficient (L/qoS)
Cy lateral-force coefficient (Y/qoS)

Cn yawing-moment coefficient (N/qoSb)

T thrust coefficient (Te/2qD2)

t torque coefficient (Q/2qoD)

L force along Z-axis; positive when acting upward

Y force along Y-axis; positive when acting to the
right

T rrnoment about Z-axis; positive when it tends to
turn nose to right

Te effective propeller thrust XR X')

YV resultant force along X-axis with propeller
operating

X' force along X-axis, propeller removed

Q propeller torque


propeller diameter (13.08 ft)










NACA ARR No. L5"09


S wing area (554 sq ft)

St vertical-tail area (19.0 sq ft as defined in text)
I distance from airplane center of gravity to
quarter-chord onint of mean vertical-tail
chord, measured parallel to fuselage refer-
ence line (19.5 ft)

b wilg span (42.83 ft)

bt span of vertical tail surface (4.25 ft as defined
in text)

ct section chord of vertical tail
q angle of yaw, degrees; positive with left wing
for rdk

a angle of attack of fuselae reference line relative
to free-stream direction, degrees

6f angle of flap deflection, degrees

6r anrle of rudder deflection, degrees; positive v'hen
trailing edge of rudder is moved to left

P propeller blade angle at 0.75 radius or angle of
sideslip,degrees

a sidewash angle, degrees; positive when flow is
from right to left r.,hen airplane is viewed
from rear

Cav average sidewash angle along rUdder hinge line
weighted for chord anr- 3-namic pressure, degrees
So1 obt
av (9/qo)a S c o

Sav
d--- rate of change of average sidewash angle with
Eang.:le of yaw

q local dynamic pressure

qo fjree-stream dynamic pressure










NACA ARR No. L5H09


ratio of local dynamic pressure to free-stream
dynamic pressure

average rlyna.c-ores'sure ratio along rudder
hinge line weighted for chord


q/qoav= Ot c dbt)


indicated airspeed


Cn*


d t

dCn
d6n

(5 r) Cn=O


(Y)CC =0
n C^


rate of change of Cn with respect to ', per
degree

rate of change of Cy with respect to i, per
degree

rate of change of vertical-tail normal-force
coefficient -.ith angle of attack, per degree


rate of change of Cn with respect to 6p, per
degree

rudder deflection at zero yawing-moment coef-
ficient, degrees

lateral-force coefficient at zero yawing-
moment coefficient


Subscripts:


vertical tail

propeller

slipstream


av average


q/qo


(q/qo)av










UACA ARR No. L5H09


AIRPLANE AND APPARATUS


Tests were made of the Grumman XF6F-4, which is a low
midwin6 single-place fighter airplane weiglhihg about
11,400 pounds and equipped with a Pratt & 1whitney R-2800-27
engine rated at 1600 horsepower at 2400 rpm at an altitude
of 5700 feet. The rear portion of the fuselage is wedge
shaped, and the gap between the rudder and fin is sealed.
The maximanm rudder travel is 3550. A three-view drawing
showing the principal dimensions and areas of the airplane
is given in figure 1 and photographs of the airplane
mounted in the Langley full-scale tunnel are given in
figure 2.

For some of the tests, the vertical tail was removed
and the gap left by its removal was faired to the contour
of the fusel~g- by a sheet of aluminum. A sketch showing
the tail fairing superimposed on the vertical tail surface
is given in fire 5, which shows also the principal
dimensions of the vertical tail surface.

The air-flow measurements were obtained by means of
the *combined yaw, pitch, and pitot-static tube shown in
detail in figure l. Photographs of this instrument
mounted in position for the air-flow measurements are
given in figure 5.


METH ODS AND TESTS


All the tests were made -'ith the &irplane landing
gear retracted and the cowling flaps closed at a tunnel
airspeed of approximately 60 riles per hour, which corre-
sponds to a Reynolds number of sOproximately .,380,000
b-sei on a mean wing chord of 7.CO feet. The ailerons
and elevators were locked at 0O deflection for all the
tests and the landing flaps were locked at 500 when
deflected. No atteirpt was made to duplicate the "blow-
up" characteristics of the lanti-., flaps. The directional
stability and trim characteristics of the airplane were
obtained for the eight representative flight conditions
outlined in table I.

Directional-stability measurements.- The directional
stability characteristics of the airplane, for each flight
condition, were investigated by measuring the forces and
moments on the airplane at :rproximat-ly 5' increments










NACA ARR No. L5939


of angle of yaw between 150, which was the maximum yaw-
angle range possible with the present airplane-support
setup in the L- ngley full-scale tunnel. For each of the
eight conditions, tests were made of the airplane with
the propeller both removed and operating and with the
vertical tail surface both removed and in place.

Directional-trim measurements.- The directional trim
characteristics c. tne airplane were determined from
rudder-effectiveness tests. Only four of the conditions
listed in table I were investigated; namely, the landing,
the wave-off, the gliding, and the low-speed climb
(Vi = 98 mph) conditions. Eudder-effectiveness tests
also were made for similar conditions with the propeller
removed.

Air-flow measurements.- Surveys of the velocity and
angularity of tn-:- '-ir lo.' in the region of the vertical
tail were made for all the conditions listed in table I.
At each angle of attack, surveys were irade for propeller-
removed and pro-eller-e-orating conditions at angles of
yaw of approxirrately 00, 5 10, and 150. The surveys
were made with the vertical tail surface replaced by the
tail fairing and consisted of measurements taken every
6 inches along a line coincident with the rudder hinge
line and extending from approximately 1i. inches above the
tail fairing to approximately 12 inches above the top of
the vertical tail surface. (See fig. 3.)

Power-on tests.- For the power-on tests, it was
desired to simulate the variations shown in figure 6 of
thrust and torque coefficient with lift coefficient for
constant-power operation at sea level. It was found that
these relationships could very nearly be produced with
a constant propeller-blade-angle setting of 24.80 measured
at the 0.75 radius; hence this blade-angle setting was
used for all the tests with the orooeller operating. A
coroarison of the variation of thrust coefficient with
torque coefficient for constant-power operation and for
the oropeller with a blade-an&le setting of 2L.80 measured
at the 0.75 radius is shown in figure 7. For the idling-
power conditions, the -ngine was run at the lowest speed
considered possible (700 rpm) without fouling the engine
spark plus. The thrust 3cn torque coefficients thus
obtained for the idling&-power conditions were 0.01 and
O.035, respectively.










NACA ARR No. L5H09


Accuracy of test results.- The accuracy of the results
of the for' tests is shown by the scatter of the test
points. The accuracy of the combined yaw, pitch, and
pitot-static tube is estimated to be about t0.25 for the
yaw- and pitch-angle measurements and about O.Olqo for
the dynamic-pressure measurements. Deviations of the
test results froni zero for apparently symmetrical condi-
tions are probably due to differences in the airplane on
the two sides of the plane of symmetry and to asymmetries
in the tunnel flow.


RESULTS AND DISCUSSION


The data are given in standard nondimensional-
coefficient form with respect to the stability axes and
the center-of-gravity location shown in figure 1. The
stability axes are a system of axes having their origin
at the center of gravity and in which the Z-axis is in
the plane of symmetry and perpendicular to the relative
wind, the X-axis is in the plane of symmetry and peroen-
dicular to the Z-axis, and the Y-axis is perpendicular
to the plane of symmetry.

The presentation of the test results and the analysis
of the data have been grouped into two main sections. The
first section gives results showing the directional
stability characteristics of the complete airplane for
the various flight conditions investigated and an analysis
of the effects of the wing-fuselage combination, the
vertical tail, and the propeller on the airplane direc-
tional stability. The results of the air-flow measure-
ments in the region of the vertical tail also are included
in this section. The second section presents rudder-
effectiveness data from which the directional trim char-
acteristics of the airplane have been determined.


DI FCTIOTAL STAbTLITY


The results of the force tests made to determine the
directional stability characteristics of the airplane for
each of the eight test conditions listed in table I are
given in figure 8. Each part of figure 6 shows curves
of Cn and Cy against for one specific flight










NACA ARR No. L5H09


attitude for the complete airplane, for the airplane with
the propeller removed, for the airplane with the vertical
tail removed, and for the airplane with both the propeller
and the vertical tail removed. No test points are shown
in figure 8 for the propeller-removed data, inasmuch as
these'data were obtained from faired curves. Values
of Cn and Cy for the complete airplane in each
flight attitude inves-tigated are given in table I.

Before a detailed discussion-is -presented of the
various factors-that affect the directional stability
characteristics of the airplane, a few of the outstanding
trends indicated by the test results-of.figure 8 are
listed as follows:

(1) The directional-stability -parameter Cy at
small angles of yaw (between t50) is smallest for the
gliding condition with flaps.retracted. For this con-
dition, C0 = -0.00015.

(2) The directional-stability parameter, at small
angles of yaw, is largest for the-high-power condition
with flaps deflected (wave-off .condition). For this
condition, Con = -0.00147.

(3) For the conditions -with-high thrust coefficients,
the directional stability decreases at angles of yaw
greater than apDroximately 100 and..inc-reases at negative
-angles of yaw greater than approximately -10.

(L) 'Flap-deflection tends -to-increase- the airplane.-
directional stability.


Effects of "'Fing-Fuselage Combination and Vertical

Tail with Propeller Removed

iing-fuselage 2or.bination.- Values of Cn, and Cy*
.for the wing-fuselage combination- are shown plotted in
figure 9 as a function of angle of attack for flaps
retracted and.flaps deflected 500. These values of Cn
and Cy, were obtained from the results shown in fig-
ure 8 for the airplane with the propeller and the vertical
tail removed. The variation of yawing-moment coefficient










'ACA ARR No. L5H09


with ingle of yaw of the wing-fuselage combination with
flaps retracted is unstable for the angle-of-attack range
investigated. Increesirg the angle of attack, however,
decreases the unstable yawing-moment variation of the
wing-fuselage combination. A further decrease in the
unstable yawing-moment variation occurs with flap deflec-
tion and causes the wing-fuselage combination to become
stable at angles of attack greater than about 80. This
increase in stability with increasing angle of attack
and flap deflection is probably due partly to an increase
in directional stability of the wing alone with increasing
angle of attack (fig. 8 of reference 5) and partly to an
increase in the directional stability caused by a favorable
effect of the wing-fuselage interference (figs. 4 and 5
of reference 4).

The variation of lateral-force coefficient with angle
of yaw for the wing-fuselage combination is positive for
the range of an.le of attack and flap deflection investi-
gated. Increasing the Pngle of attack and deflecting the
flnps decreases the rate of change of lateral-force coef-
ficient with angle of yaw.

Air-flow surveys.- The results of the air-flow
measurements for the propeller-removed conditions are
given in figure 10, which shows the variation with
height above the fuselage a-long the rudder hinge line
of the sidewash angle o and the dynamic c-pressure
ratio q/qo for angles of yaw of approximately 00, *50,
100, and 150. Weighted average values of the sidewash
angle and dynamic-pressure ratio along the rudder hinge
line are given in table II.

The surveys (fig. 10) show that, for this airplane,
the variation of average sidewash angle at the vertical
tail with arigle of yaw. do/di} was, in general, positive
(destabilizing). The data show that the direction of
flow from the fuselage wake and air beside it (region in
which sharp loss in dynamic pressure occurs) is strongly
destabilizing. Inasmuch as the vertical-tail chord is
largest near the fuselage, the effect of the flow in this
region on the contribution of the vertical tail to the
airplane directional stability should predominate. The
flow above the fuselage wake appears, in most cases, to
be slightly destabilizing for negative angles of yaw and
to have little effect on the stability at positive angles
of yaw. Increasing the angle of attack or deflecting the
flaps tends to increase the destabiliz'ing effect of the










NACA ARE No. L5H09


sidewash. These results are, in general, contrary to the
results published in reference 5, which indicate that the
sidewash is usually stabilizing for lov:-wing airplanes.
The discrepancy may be due to the fact that, for the
present series of tests, the horizontal tail and canopy
were in place and the rear portion of the fuselage was
wedge shaped; whereas the tests of reference 5 were made
on a smooth circular fuselage with no horizontal tail.

The data given in table III show that the dynamic-
pressure ratio at the vertical tail has its minimum value
at small angles of yaw and increases as the angle of yaw
is increased in either direction. For any given angle
of yaw, the contribution of the vertical tail to the air-
plane directional stability is directly proportional to
the dynamic-pressure ratio at that angle of yaw. At small
angles of yaw (between 50) the vertical tail lies directly
in the path of the fuselage and canopy wakes and hence
q/q for these conditions reaches its minimum value.
As the angle .of yaw is increased in either direction, the
vertical tail moves away from the fuselage and canopy
wakes and q/q. increases. Inasmuch as the fuselage
boundary-layer and canopy wakes increase with-increasirng
angle of attack, the loss in q/qo at the tail increases
with increasing angle ..of attack.

Vertical tail.- Ex.perimental increments of yawing-
moment and lateral-force coefficients due to the vertical
tail were obtained from the data of figure 8 for the
prooeller-removed conditions and are shown plotted in
figures 11 and 12 for all the airplane attitudes investi-
gated. Figures 11 and 12 show also increments of yawing-
moment and lateral-force coefficients due to the vertical
tail that were computed on the basis of the results of
the air-flow surveys.

The force-test data show that the contribution of
the vertical tail to the airplane directional stability
is.lower in the yaw-angle range between -50 and 50 than
at the higher angles of yaw and, in addition, the contri-
bution of the vertical tail decreases with increasing
angle of attack and flap deflection. Numerical values
for the slopes Cnt and C are given in table III.

The trends. shown bythese results are in agreement with
the conclusions drawn from the results of the air-flow
surveys.










ITACA ARR 'To. L5HO9


An analysis has been made of the results of the air-
flow surveys and the force tests in order to investigate
methods for co.nputing the contribution of the vertical
tail to the airplane directional stability. The incre-
ments of yawing-moment and lateral-force coefficients due
to the vertical tail are given by the following expres-
si ons:

S \av)(q/qo) St

ACnt T o Sb (1)



= (2)
ACyt ACnt (2)


The values of oav and (q/qo) v in equation(1), which
were determined from the air-flow surveys, are assumed
to apply to the small area below the lower limit of the
air-flow measurements.

The results of the air-flow surveys when used in
conjunction with the recommendations given in reference 1
with re-ard to the deterinination of the tail area, tail
scon, and tail lift-curve slope were found to give
values of Cnt and CYt that averaged about 20 percent
larger than the values obtained by the force tests. The
values of the vertical-tail ers. and vertical-tail span
determined by the methods of reference 1, however, include
areas in excess of that part of the vertical tail above
the fuselage. The surveys inic.cated that the contribu-
tion of these areas to the airplane directional stability
would be small because of the large destabilizing side-
wash and low dynrror'ic pressure in that r.-,-'on. Conse-
quently, for further calculations, the area of the vertical
tail was considered equal to the actual vertical-tail area
removed frorr the airplane during, the tests (St = 19 sq ft)
and the span of the verti-al tail was considered equal to
the height of the vertical tail above the too of the tail
fairing (bt = 4.25 ft). (See fig. 3.) All the terms of
equation (1) except (dCT/da) are Vnown from either
the surveys or the force tests. The term (dCN/da)
includes the end-plate effect of the hori-ontal tail and
fuselage on the vertical tail referencess 1 and 6)










NACA ARR No. L5H09


modified by the interference effect of the vertical tail
on the fuselage. The lift-curve slope for an isolated
tail may be determined from figure 5 of reference 1 as
a function of tail aspect ratio. The analysis of the
results of the force tests and the air-flow surveys
revealed that the geometric aspect ratio of the vertical
tail bt2/St should by multiplied by 1.55 to account for
the end-olate and interference effects. Although this
value is numerically the same as that recommended in
reference. 1, the agreement is coincidental in view of the
difference in definitions of tail area and tail span. The
comparison given in figures 11 and 12 of the increments
of Cn and Cy due to the vertical tail, as determined
from the force tests and as calculated from equations (1)
and (2) by use of the air-flow-survey data and the correc-
tion factor of 1.55 for the geometric aspect ratio of the
vertical tail, is given to show the range of application
of the present method for the XF6F-4 airplane. Good
agreement is obtained for the complete range of angle of
attack and yaw for all conditions investigated.

In order to calculate the contribution of the vertical
tail to the airplane directional stability, the variation
of sidewash angle and dynamic-pressure ratio with angle
of yaw n'ust be known because


C d(i a )(q/qo) St L
CnGt d t d* S b

and

y = -nt (4)
t It "

Equation (3) shows that the contribution of the vertical
tail to the airplane directional stability is directly
proportional to the derivative of ( cav) (q/qo)av with
respect to the angle of yaw. The term (V oa0 )q/qo)v,
which is designated the air-flow factor, is shown plotted
in figure 15, and average values of the slopes
d "' -a)q/qo)av between = -5 and = 5 are
di i
given in table III. This table indicates also the effect










NACA ARR No. L5H09


on the contribution of the vertical tail to the airplane
directional stability of the decrease in the derivative
of the air-flow factor with angle of attack and flap
deflection. For test conr'itions with flaps deflected 500,
the destabilizing effect of the sidewash cnd the loss
in q/qo is sufficient to reduce the contribution of
the vertical tail to the airplane directional stability
by about 50 percent of the value that would be obtained
Sd(^ a T q/q
d av oav
if were equal to 1.0. The com-
dV
prison given in table III of the values of Cnt and

C obtained front the force tests and calculated from
t
equations (3) and (L) by use of the air-flow-survey data
and the correction factor of 1.55 for the geometric tail
aspect ratio shows fairly good agreement between these
slopes.


Effects of Propeller Operation

7ti- total increments of yawing-moment and lateral-
force coefficients due to propeller operation are given
in figure 1k for each of the conditions investigated.
These increments were obtained from the experimental data
plotted in figure 8 and are the differences in Cn
and Cy for the complete airplane r'th the propeller
operating and the propeller removed.

For the airplane with f.r-ps retracted (fig. l1(a)),
propellor operation was destabilizing at angles
of yaw from about -100 to 15) t-;" instability was
gr -test at l'?rF- positive angles of yaw. At arnles of
*r between -100 and -150, p-oreller o-j-ration gave a
stable variation of ACno against V. Pone of the
effects of propeller operation was proportional to the
oower appli-,. or to the thrust coefficient; in fact, at
small angles of yaw (between -50), the instability
caused '..- propeller operation was about the same for
all conditions, regardless of the thrust coefficient and
angle of attack. The effect of propeller ooer'ition on
the directional stability of the airplane with flaos
deflected 500 at small angles of yaw (fig. lh(b)) was,
in general, to increase the stability for the wave-off
)o1:2 tion, to decrease the stability for the lfr:,i :-









NACA ARR No. L5HO9


condition slightly, and to cause no appreciable change
in the stability for the landing-approach condition. The
average increase in directorial stability due to propeller
operation for the wavo-off condition (rated power,
Tc = 0.51), at angles of yaw between 50, was very large
Scn = -0.00105).


The effects of oropeller operation on the directional
stability characteristics of the airplane can be con-
veniently considered under the following groups;

(1) Direct effect of the propeller forces on
the airplane directional stability

(2) Effects of the propeller slipstream on the
contribution of the wing-fuselage combination to the
airplane directional stability

(3) Effects of the propeller slipstream on the
contribution of the vertical tail to the airplane
directional stability

Direct effect of propeller forces.- Methods for
computing the direct effect of the propeller forces on
the variation of lateral-force and yawing-moment coef-
ficient with angle of yaw are given in reference 7. The
dashed lines shovm in figures 15 and 16 are increments
of C and Cy due to the propeller forces that were
calculated by equation (7) of reference 7. (The pro-
peller side-force factor was 99.2.) The calculations
show that the direct effect of the propeller forces is
to decrease the airplane directional stability for all
conditions investigated. This effect is greatest for the
low-speed climb condition (CL = 1.59, T = 0.51 for
which the decrease in directional stability due to the
isolated propeller is 0.00058.

Effect of slipstream on wing-fuselage combination.-
The effects of the propeller slipstream on the lateral-
force and yawing-moment variations with angle of yaw of
the wing-fuselage combination may be indirectly obtained
from the experimental results. The increments of Cn
and Cy due to propeller operation for the airplane with
vertical tail removed, increments which were obtained
from the force tests, are shown by the solid lines in









NACA ARR No. L5HOq


figures 15 and 16 for each condition investigated. These
increments include the direct effect of the propeller
forces and the effects of the passage of the slipstream
over the wing-fuselage combination. The difference between
the solid and the dashed lines in figures 15 and 16 are
therefore presumed to be due only to the effects of the
slipstream on the wing-fuseleae combination.

The data show that for all conditions with the flaps
retracted, at angles of yaw between 50, the slipstream
effects on the wing-fuselage combination caused destabi-
lizing variations of yawing-moment coefficient with angle
of yaw. At the low thrust coefficients this effect was
small; at Tc = 0.51, however, the slipstream caused a
destabilizing increment of Cn of about 0.00047. For

the flaps-deflected conditions, the directional stability
of the airplane was not changed appreciably by the slip-
stream effects on the ving-fuselage combination for angles
of yaw between 5 and -15 but was considerably decreased
for angles of yaw between 5 and 150

Effect of slipstream on air flow in region of vertical
tail.- The results of the surveys with the propeller
operating are given in figures 17(a) to 17(e) for the
flaps-retracted conditions and in figures 17(f) to 17(h)
for the conditions with flaps deflected 500. Weighted
average values of the sidewash angles and the dynamic-
pressure ratios at the vertical tail determined from
these surveys are given in table IV.

For all conditions investigated, the variation of
the average sidewash angle at the vertical tail with
angle of yaw was generally destabilizing (positive doav/d .
The destabilizing effect of the sidewash appeared to
increase with thrust coefficient and angle of attack and
to decrease with flap -eflection. (See table IV.) The
most important factor contributing to the destabilizing
effect of the sidewash is the flow from the fuselage
boundary layer, which exists in the region in which,
for the present airpli-ne, the vertical-tail chord is
largest. The destabilizing sidewvsh in the region of
the fuselage boundary layer was smaller in magnitude for
the flaps-deflected conditions (figs. 17(f) to 17(h))
than for the fl-rs-retracted conditions (figs. 17(a)
to 17(e)). The data show that the air~ flow -t the vertical
tail in the region above the fuselage boundary layer is









NACA ARR No. L5H09


dependent on the conditions of propeller operation. As
the thrust coefficient increased from one condition to
another, the sidewash in this region became increasingly
negative (flow fror. left to right when airplane is viewed
from the rear). This effect may be accounted for by the
slipstream rotation. The vertical tail was in the region
of the rotating flow from. the uoper half of the propeller,
which for right-hand propeller operation caused the air
to flow from left to right. A further effect of the pro-
peller rotation was a lateral displacement (toward the
right) of the slipstream in the region of the vertical
tail due to the tangential-velocity components of the
rotating flow. The result was that,as the airplane was
yaved nose left (negative yaw), the vertical tail tended
to move into the center of the slipstream and the side-
wash become increasingly negative; as the airplane was
yawed nose right, however, the vertical tail tended to
move awa y from the center of the slipstream and the side-
wash became decreasingly negative. These tendencies
indicate that increasing the slipstream rotation tends
to increase the destabilizing effect of the sidewash.

The effect of the increased dynamic pressure at the
vertical tail due to the propeller slipstream was to
increase the contribution of the vertical tail to the
airplane directional stability, inasmuch as the average
sidewash was never large enough to cause the contribution
of the vertical tail to be destabilizing. Surveys
(fig. 17) shoved that the disposition of the slipstream
at the vertical tail was such that the maximum dynamic
pressure occurred at the sections near the middle of the
tall for zer. angle of yaw and at the sections about one-
third of the tail height above the top of the fuselage
for other angles of yaw. The dynamic pressure was a
minimum at the bottom of the vertical tail as a result
of the large dynamic-pressure losses due to the fuselage
and canopy wakes. The displacement of the slipstream
with respect to the vertical tail, as the angle of yaw
is changed in either direction, can be observed from the
dynamic-pressure maasurm:--nts. The results (fig. 17
and table IV) show that the dynamic pressure at the
vertical tail is highest for negative angles of yaw
and is lowest for positive angles of yaw. These results
indicate that the contribution of the vertical tail to
the directional stability of the airplane with the propeller
operating will be greatest at negative angles of yaw.

Effect of slipstream on vertical tail.- Experimental
increments of lateral-force and y&wing-moment coefficients









NACA ARR THo. L5H09


due only to the effects of the oropeller slipstream on
the vertical tail surface were obtained from the data of
figure 8. The increments, which are the difference between
the increments of Cy and OC due to the vertical tail
with the propeller operating and with the propeller
removed, are shown in figure 18. In general, these
results substantiate the conclusions drawn from the air-
flow surveys in regard to the effects of the propeller
slipstream on the vertical-tail contribution to the air-
plane directional stability. The variation of ACnt
na
with angle of yaw is such as to decrease the airplane
directional stability at high positive angles of yaw
and to increase the directional stability at high negative
angles of yaw. Except at Tc = 0.01, at which the
effects of the slipstream are small, the directional
stability is increased for all conditions in the low-
yaw-angle range (between 50) as a result of the slip-
stream. This stabilizing effect of the slipstream at small
angles of yaw increases as the thrust coefficient increases.

The total increments of Cn and Cy due to the
vertical tail are given in figures 19 and 20 for the con-
ditions with the propeller operating. These increments
were obtained from the data of figure 8 as the differences
between the propeller-operating results with the vertical
tail installed on the airplane and with the vertical tail
removed. Also shown in figures 19 and 20 are increments
of Cnt and CYt that were calculated from equations(l)
and(2)by use of the air-flow-survey data with the pro-
peller operating and the effective lift-curve slope of
the vertical tail determined from the data for the
propeller-removed conditions. Curves showing the varia-
tions of the air-flow factor with angle of yaw for the
propeller-operating conditions are given in figure 21.
The agreement between the calculated and the experimental
results shown in figures 19 and 20 is good.
d(ii oav)(q/qo)av
E:perimental values of the slope -,
which is used in equations (3) and (4) for calculating
the contribution of the vertical tail to the airplane
directional stability, are given in table V. These
values show that the effect of the vertical tail in
increasing the airplane directional stability is greatest
for the conditions with the highest thrust coefficients
and decreases as the thrust coefficient decreases.









NACA ARR No. L5H09


numericall values of C and Cyt obtained from the
t t
force tests and calculated from equations (3) and (4)
by use of the air-flow-survey data and the tail lift-
curve slope previously determined are also given in
table V. The satisfactory agreement between the results
indicates that little change in the effective slope of
the lift curve of the tail occurs as a result of the
propeller slipstream.


DIRECTIONAL TRIM


The results of the rudder-effectiveness tests are
given in figures 22(a) to 22(c) for the airplane with
the flaps retracted and the propeller operating to
simulate a gliding condition and two low-speed climb
conditions and in figures 22(d) and 22(e) for the air-
plane -:ith the flaps deflected 500 and the propeller
operating to simulate a landing and a wave-off condi-
tion. The results of the tests with the propeller
removed are given in figure 23 for the airplane with
flaps retracted and with flaps deflected 500. The more
important results of the rudder-effectiveness tests are
summarized in figure 2L, which shows curves of dCn/d6r,
6r)Cn0, and (C)C,=0 plotted against angle of yaw

for each condition investigated. All the values of the
slope dCn/d6r were measured at zero rudder deflection
as a basis for comparison.

For the propeller-removed conditions, dCn/d6r
reaches its minimum value near zero angle of yaw and
increases as the angle of yaw is increased in either
direction (fig. 24). The dynamic-pressure losses at
the vertical tail are greatest at zero yaw, and the
losses decrease as the angle of yaw is increased in
either direction. For the propeller-operating conditions,
the rudder effectiveness increases as the thrust coef-
ficient increases from one particular condition to
another because of an increase in the dynamic-pressure
ratio at the vertical tail (fig. 24). For all the con-
ditions investigated with the propeller operating,
except the gliding condition with flaps retracted,
dCn/d6r attains its rraximum value at high negative
angles of yaw and its minimum value at high positive









NACA ARR No. L5H09


angles of yaw (fig. 24); the dynamic pressure at the
vertical tail reaches its maximum value for high negative
angles of yaw and reaches its minimum value for high
positive angles of yaw. An analysis of the test results
showed that the values of dCn/d6r are very nearly
directly proportional to the dynamic-pressure ratio at
the vertical tail.

The rudder deflections and angles of sideslip required
to trim simultaneously the airplane yawing moments and
lateral forces for each condition investigated were
determined from the data of figure 24 and are given in
table VI. For the conditions with the propeller removed,
the data show that the values of 6r and p for zero
yawing-moment coefficient are small. For the conditions
with the propeller operating, the data show that the
rudder deflections required for directional trim are
greatest for the two low-speed high-power conditions.
(See table VI.) These deflections, however, are con-
siderably lower than the maximum available rudder travel
on the Grumman XF6F-4 airplane.

The data show that the amount of rudder deflection
required for directional trim in any condition is
primarily dependent on the effects of the propeller
slipstream on the vertical tail and on the wing-fuselage
combination and, to a lesser degree, on the direct effect
of the propeller forces. The increments of Cn and Cy
at zero yaw due to the effects of the slipstream on the
vertical tail, the effects of the slipstream on the wing-
fuselage combination, and the direct effect of the pro-
peller forces are given in table VII for the wave-off
and low-speed climb conditions. Of the total increment
of Cn at zero yaw due to propeller operation for the
low-speed-climb condition, 77 percent was due to slipstream
effects and 25 percent was due to the effects of the pro-
peller forces. For the wave-off condition, 98 percent of
the total increment of Cn at zero yaw due to propeller
operation was caused by slipstream effects.

The curves in figure 24 of (Sr)C =0 against ,
besides indicating the rudder deflections required to
trim the airplane yawing moments, are a measure of: the
airplan- directional stability. The conclusions
regarding the airplane directional stability character-
istics, which are derived from these results, are sub-
stantially the same as those derived from the curves of











figure '. showing the variations of Cp a-sain3t
for ,5 = 0.


SU 7. MMAR Y OF RES ULTS


Data are presented of measurements nade in the
L,'n-ley full-scale tunnel on the Grumman XF6F-4 airplane
to investiAte the factors affecting the directional
sti-ilitjy ,nd tr5-. characteristics of a typical fi-hter-
t;-,e .irplane. Although these data are quantitative
for this particular airplane, the t.rnds are believed. to
be ener-illy applicable to reasonably similar airplanes.
The resultL are su-rmmrized as follows:

1. For the conditions investigated, the value of
the directicnal-stability parameter Cn at angles of
a ,,' between 50 ras smallest for the gliding condition
ith llaps retracted (O, = -0.00015) and was largest
for the ..vi'.e-off condition with flaps deflected 500
I'" = -0C'1l?7). With the values measured in the
low-?,v'-an.le range used as a reference, the airplane
directi.,:-.:l stability for the conditions with high
thru3ust co,.fficients was decreased at lar.-e positive
er!-:ls of :aw arnic was increased at large negative
an .'-- cf -:aw.

.', the XF6F-64. airplane, the variation of aver se
si :!e' a-i --'-_gle at the vertical tail with angle of yaw
was s--ri.':.ily uch as to decrease the contribution of
the vertical tail to the airplane. directional stability.
Prop)eler operation increased the ,n -r.i.tude of the
destalili?.r3: effect of the s.idewash but, at small 1a rls
of ;:r, a.o increased the dynamic pressure at the tail
suff:ci.*~?ntly to make the combined effect stabilizing.

5, fri- wing-fuselage combination with flaps retracted
was d3 rectionally unstable for the ran-le-of-attack rn-. e-e
investi~.ted. Increasing the single of attack aj.d
deflectiniL the flaps decreased the unstable variation of
ya-',;n.-.T'o-.,nt coefficient with asn-le of yaw of the wing-
fus-':s::e combination.

1.,. For all the conditions investigated with the
fla- retr-cted, the contribution of the propeller
dj.creza.ed th directional stability of the airplane at
small angles of yaw. 'Vith the flr-is deflect-.d 50O at


F'10A ARR iTr. L J!'D9









NACA ARR No. L5HO9


small angles of yaw, the contribution of the propeller
increased the airplane directional stability appreciably
for the wave-off condition, decreased the airplane
directional stability slightly for the landing condi-
tion, and caused no appreciable change in the stability
for the landing-approach condition.

5. The propeller slipstream increased the contri-
bution of the vertical tail to the airplane directional
stability at small angles of yaw. As a result of the
lateral displacement of the slipstream with respect to
the vertical tail, the contribution of the vertical tail
to the airplane directional stability was greatest at nega-
tive angles of yaw and was smallest at -positive angles of yaw.

6. The destabilizing contribution of the wing-
fuselage combination to the directional stability of the
airplane for the cor:ditions with the flaps retracted, at
angles of yaw between i50, was increased by the effects
of the propeller slipstream. The directional stability
of the airplane for the conditions with the flaps
deflected 500 was not chcn.ed appreciably by the slip-
stream effects on the vrinp-fuselage combination at angles
of yaw between 50 and -15 but was considerably decreased
at angles of yaw between 50 and 150.

7. The amount of rudder deflection required for
directional trim is primarily dependent on the slip-
stream effects and only secondarily dependent on the
direct effect of the propeller forces. Of the total
increment of yswvng-moment coefficient at zero yaw due
to propeller operation for the low-speed climb condition,
77 percent was due to slipstream effects and 25 percent
was due to the effects of the propeller forces. For
the wave-off condition, 95 percent of the total increr.ent
of yawing-moment coefficient at zero yaw. due to propeller
operation was caused by slipstream effects. The wave-off
cornditior, at a lift coefficient of 1.39, required the
largest amount of rudder deflection for trim (6r = -18.50)

8. A comparison of the results of the extensive airflow
surveys with the results of the force tests r-'e possible the
determination of a value for the effective-lift-curve
slope of the vertical tail; this value' permitted







NACA ARR Fo. L5H3?


calculation of the contribution of the veitictl tail to
the directlnsfl stability of the airplane within accet-
able limits.

Langley Memorial Aeronautical Laboratory
National Advisory Comirttee for Aeronautics
L&nrley Field, Va.





RE FE RENCES


1. Pass, H. R.: Analysis of Wind-Tunnel Data on Direc-
tional Stability and Control. lACA TN !o. 775, 19L-0.

2. Imlay, Frederick H.: The Estimation of the Rate of
Ch'-ine of Y i.ng :'nome.-nt with Sideslip. NACA TI
No. 656, 1938.

3. Shortal, Joseph A.: Effect of Tip Shape and Dihedral
on Lateral-Stability Characteristics. 'T.CA
Rep. i:. 548, 1955.

h. Recant, I. G., and Wallace, Arthur R.: Wind-Tunnel
Investigation of Effect of Yaw on Lateral-Stability
Characteristics. IV S- r 1?trically Tapered "ing
with a Circular Fuselage Having a "'edge-Shaped
Rear and a Vertical Tail. NACA ARR, March 19.2.

5. Recant, Isidore G., and ''allace, Arthur R.: Wind-
Tu.:nel Investigation of the Effect of Vertical
Position of the .ling on the Side Flow in the
Pepion of the Vertical Tail. NACA TI No. CrL,
191l.

6. Katzoff, S., and IYutterperl, ""illiam: The End-Plate
Effect of a Horizontal-Tail Surface on a Vertical-
Tail Surface. TA A TN No. 797, 1911.

7. FRbner, Herbert S.: !Jotes on the Propeller and
Sliostream in Relation to Stability. iLA:A ARR
No. Lhil2a, 19LL.












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0 o H H Ho

bo0 -d O n co Xt J O CaOj
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O rl K\ C-4 r-
r- r-


ClO
0


S -.
oo

, E-,
O 0





Z-4

0 '
04
OO










NACA ARR No. L5H09


OO
-o NJ N 0 1 O 0 t) CO _
S00 .



0 U W- 1
,- O --O 2j 0 N 0 O

(0 U) O l \i rH L N N\ Ct K\
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So U rH CO N o -j- H




S 10 0 o0 0o 0 0 0 0 0






E 00; 0
I @ F, i II II

0 E-4 E-a 0 E-4 0
0
H 0 0 0
) C II aI II


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0_ V1 0 X0E E-
O MP P 4
9 oa ca o ca o


I I I I
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I I I H H H I
I I I I 0 0 C'
* I I I
I I I I


o



coo
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0
SE-









NACA ARR No.


O




~I.
0


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<






Sa


0
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0 0




SI
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<3 l

C



o
O3 (




a
0 s
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a

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00 0
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k 0o 0 O O











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0 4- <. *








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oiI o a o o
doi .o o *


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Il II iI
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ae a
as a


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C II II I II
o0 0 II 0 -1
U Co H > 0 0 0
I0o d -

_______>} _______O


0



Oo

00




OH
HM







NACA ARR No. L5H09


- : : 4s ;



: -:


a ,- t



t* t3..


*. t3






?g^


.o-

<:z





^t
1



.- x






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L4JQ Z Q
hasat
444 44


Fig. 1


*
.9-
P0



Lk



q,






a
q)




I

9-
0
*S*
F








NACA ARR No. L5H09


Fig. 2a








NACA ARR No. L5H09


Fig. 2b





NACA ARR No. L5H09


"a


Luj


Fig. 3


1



~I







NACA ARR No. L5H09


-Ir
(5


0(


aa
on


ZuJ
I-8

o
U<


1-
0
4-,





C3


,Z)
0
-`2


Ij,




-'zz
0


O


Fig. 4


a 4% >\
AX








NACA ARR No. L5H09 Fig. 5a














CD





r-i
44
0.0
3 (4a



bo
i .,
I i: ..


D 0.
a:'* ,






d a













f-ii
d El. L
:: .








NACA ARR No. L5H09


Fig. 5b








NACA ARR No. L5H09


Horsepower En9e rpm
Rated power /600 2400
--- o.e rafed power /040 /960








.6


.4


.2
NATIONAL ADVISORY
--- COMMITTEE FOD AROIAUTICS
11


.2 .4 .6 .8 /0 /2
L/ft coeff/1/ent CL


/4 /.6


/8 20 22


Figure 6 Calcu/ofed var/'r/bns of Tc and Qc
with CL for con stant-power operation
at sea /eve/.


0


Fig. 6







NACA ARR No. L5H09


Horsepower


Rated power /600

- 0.65 rated power /040

------- 9, 4.8 measured at 0.75 R


/ :


2-.


.I -


'3



UP


.C


Rated
I I


power
UOC~


I I I I


Thrust


0.65


rated power


NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS
l I I I I -i -i --i --...


coefficient, T


Fig-re 7. Comparion of variation of T with Qc

for constant-power operoton and for


the propeller


with blade angle fixed


of 24.8 at 0.75 rodius,


Engine rpm

2400

/960


Fig. 7








NACA ARR No. L5H09


.2





0





-.2


Vertical foil Lift coefficient, C
On Rated roao 0.24
- Off Rated C=O 4
-On Prope/ler off .3
Off Propeller off .&3


S.04


t O-
02



0
Ol



S-.04


NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS
: : 11 1 1 1I I I I I I I I I I 1j


Figure 8.-


-s 0 ( 0 / .5'
Angle of yaw, y, deg
(a) C/imbing attitude; oc,/~Oy, dfot
Directional- stablhfy characteristics of
XF6F-4 airplane; rudder /ocked at
0" when on airrplane.


4 ---- 9- --0- ._

- ---- ,,,,,,_

Fig. 8a








NACA ARR No. L5H09


Vertical fail
On
Off
On
Off


Powr Lift coe~ffcientC
Rated7=o. 0.4
Ratec=O.) .43
Propeller off .4-0
Propeller off .40


.-4:


NATIONAL ADVISORY -
COMMITTEE FO AERONAUTICS


0o -s 0 s /0 15
Angle of yaw, y, ceq


(b) Climbinq attitude; oc, 3.4", 6,0.


s .04
Aj

S.02
8


o
E
&-.oz
s -4
a-..l


S-'


EEE tIW i-i


Figure 6.- Continued'


Fig. 8b


--~-







NACA ARR No. L5H09


Vert-rc/ 77//


Power L,/f coefflic/er, C,


0
-0-


Or7
Off
On
Off


Rated (=0. 0
Rated (7= 0.30
Prope//er off
Pro'el/er off


0.96
.96
.80
.80


_-0 4 1 1 I 1 1 1 1 1 1 1 1 1 1 1 1 1 1--i -- 7 ; .
-5 -/0 -5 0 5 /0 /t
Ary/e of yaw r, dey


(c) C/lmbinc affifude; oc,98.9; f,0O.
Flqure Confinued,


Fig. 8c


r








NACA ARR No. L5H09


r -


Vertical tail
S On
------ Off
--- On
-- Off


Power
,eafed (Tra ) 0.5/)
Rated (T= o.5
Propeller off
Propeller off


Lift coefficientC
1.39
1.39
1.04
/.04


-L-





NATIONAL ADVISORY
SCONNITTEE FOR AERONAUTICS

'5 -/0 -5 0 5 /0 /5
Angle of yaw, I dey


(d) Climbing atfhtude; -oc,/Z.3. f, O*.
Figure 6.-Continued.


.02





-.OZ


-.04


Fig. 8d







NACA ARR No. L5H09


Vertical ta/i
On
Off
Off


Power L/ft coeff/c~,e C,


S= 0O.O/
T = 0.0o/
Propeller off
PrOpe/Ier off


0.83


' A


- -'-- ""----- -------
O-


)2
NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS
4 1 1- '* I '* I I I I l l l 1 -


-5 0 5 /0


Ang/e of yow t,


deg


(e) Glding attitude; aoc,9.2L" f,".
Figure 8.- Continued.


p


-/5


T I !






I JAiiILA
EEEEEEEEEEEEEEEEEEEE-E;:1


.(


-


-(


Fig. 8e







NACA ARR No. L5H09


. . I -


..I : "'j Ii II: I I: I lii: I :'''''


Vertcal fail Power Lft coeffie/nt/, C,
-- On O. 'Raed(ite=O.S3) /.37
Off .65Rated ec=03_) /.37
on rope/r o /.//
--- Off Prope/ler off ///





.04



-02
.04 | ||,l-, | -_.
-^ "E: :::: :::: :: ......::::-,


NATIONAL ADVISORY
COnMITTEE FOR AERONAUTICS -


5


-/0 -5 0 5
Ang/e of yaw, ?, o'eg


(f) Landing- approach athfude; oc,~ ; sf,SO".
Figure 8.- Continued.


.0

0




-.2


-041
-I7


/0 /5


LL LJ- l I I IIII I I I I I I I I I I 11 1 11 1]


~ii~i~i~i~ii~i~


f___ L


i


,


Fig. 8f







Fig. 8g


NACA ARR No. L5H09


Vertical tail
On
Off
On
Off


Power
Rated CTc= oas5
Rated Tc-= 0./)
Propel/er off
Propel/er off


Lift coeffiient, C
1.39
".59
1.04
/.04


NATIONAL ADVISORY
COMMITTEE FOI AERONAUTICS


-/5 -10 -5 o 5 /0 IS
Angle of yaw yr, od


(g) Wave-off atfhude; oc, 4.9; df,50.
Figure 6.- Continued.


-a---


::::::::::::::::::: ;
: ^ : : : :,^ :.', 7^


_ I ^ ri.,,,^ -

k,,__"i Z Z
:^ e,=EE:"T=EE:EE








NACA ARR No. L5H09


Vertical taf/

On
Off
--- Onf
- -- Off ,


Power L;
7c = o/
e = 0. 0/
Propeler off
-rope//er off


ff coeffic/ent, C1

/.58
/.56
/.56


NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS
I I I I I F I i l ll Il


/0 -5 0 5 /0 /5
Ang/e of yaw, ~, deg


(h) Landing aHltude ; cc, ia.; ,5fso.
Figure 8.- Concluded.


.04,


-.04
-/1


5


J .- I -_ I_


Fig. 8h






NACA ARR No. L5H09


t~HZ4


/Z


/6


I I I


F/las retracted

F/ap1,s defected JO0


Z ZIZ[1111


NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS


H1 IZLZLII


SF/ap s refracted


- r t~fi17


_U_ __ __~~/elJeL5efece


4 8
Ang/e of attackr


/2 /6
, c deg


F/iure 9 Contribuf/on of win- fuselage
combination o-. Cn and Cy Vertical
fail and propeller removed.


.01


(..


-.01


0


-001


50


Fig. 9


.001/


-








NACA ARR No. L5H09 Fig. 10a


- .r -r --< ,I-'-
--i d-.-I- --- ;- _t i---" _
, ,-I ,6. T-



r-1 -




....t :4-4-4
.h .








NACA ARR No. L5H09


Fig. lOb









NACA ARR No. L5H09


'" 'rU:4. .'> >. -ff44i
-: -t ,-'-l:j .-i '-
,--. .. T sh ;- *


tHi:dz s LJ Dt-IAi"


N- 4' i i-,.L 2 --i-I- -h I i. *. '4-


'-i~~~g z- Y I- r :tt -- TT-H, ~I: C:;-; _4r -^ 7rT-s Wt--'^


--4



1
IIt






-- -----'
--------'--- F .-- -. -
^s~ tS^- : .: .*









t F F

I- j -J. Y1 ;
-!^ l--l-^ Tarl 1, TMTilrp l














LhtR
iti

-1 L 4
1;:
Fd +i- P0












"-, T- tP
N CIA I 1f

4 =0 F? i
-- ty L f,
1-" 1 T0



r ~ no 'i7 '

r-
ts~sT~i~i^^l^^^^i^sipisl^^|?^^%p


~*&~-1-*-f-~-F Nd-Htf~ PHH 4tH 4+


i- .-


I~i? i


.
i Lr L1 i"l*^

.f j -'


." 1 "1-


F-igft-'lA.


im 'lrk i -9


RIM
26

^fi^r^W -z:'

rT iF
t^ ^ ^47.


if i"'y-E~l -1."i~-t "--V it~.


~ll--l;t t
_t


Fig. 10c


7-


Il
":
.,7


m t K- -_... ...-... ,-,-,-:










NACA ARR No. L5H09


^r -I


! _-L


F FFFRF ffF


F -- ti-] i
nf


--


-t
I-. -;- H


1-X







i









I f
r F- -~







-r_- A. 1
+b
-r I' + -1
^^- ^l i'[ -1 ? It ??-1 1 A












t fi .* rfFf;i i ? f F




1
=1^ ^^f=-^ 1^^I ^ ^ F4:: ^: 1: upj+- F. F-.:

II^ ^ ,il^ -|_ ^ ( -_ --- -J- ).- p 4 .+ ^ 4 -^ --^ --- ^ t-;- -!


















;if
^ rX 1: 1 11I E ^^ I -:^ ^^



















;ii
:41'
!.3e^ -^ --- ^ -- X 4 -- -- 4 -t;?;rL-+:4 --:- i


- T-H- H- H-H+


-- 7-f-L -j,. --H _-H+H- -T-H N+ fiH


n b lst l -r twjr L V4 l-r P i-n 4 I | I


_ ~_~_l~rr~_ ___i


I.


I


T V


LrJu --t r-r-QV;


Fig. 10d











NACA ARR No. L5H09


4L^^ t ..4I^f 1'.






-4-t

fil J







3-1






rFT





Lt -L L t
-777777-
Eii l4


























3. -i -
'I i-f 1




r














F ---- :--.
t Lr

+ '-''*~~ 1^'' '^ ^ 1 1 1fT -~-~'~ ssz-' '*B"-'r '7~i =!;^E -s F. -
^^ --t^ --: ^-f+^L ^. ^ ^ -^ ^: :: ^;- ?4

ise~^~j.p .^^~fS^^^E^^;^^!



.-Li-^:^: ^:-:^ jfi'^ f^ i: ^ ^ ^!;^ ^


*^n^^i^^^^ ,::^^^:!^ ^!; ;










'|^ Ss S^E |i^ ;:u 2 :aE ^;:^:: :i

^4 ^ar '. ,ua^ -- l -- 'j7*" :* ? '- -*2:---^ *- l' h -*i'--*-


. .- F J T 4 E2. 11 L L. fA A FIE Ti III SAT


Fig. lOe









NACA ARR No. L5HO9


-.- I --- _II- L -i-i .-.i '
'- "-'


__ t-*I t#, i. -'t $ .


4ThWT~fI'HtI


. H'"- -t
_ I -
: l


zk1ifl N
ttik'
-iv it


iri
:^


t


::.1


r7


t~


--- -4-

-; -. _L


-ii

*Hc rtrnr


Th~
l~t1yLi


I-i


A-Aj -Ti* -
-Is' s^ "!
Tp 'l&lf-


I:- .l- :i _-I -i I- ~ -~~-i- m 1t--.-

i- t-1 .J .. .,- ...I- J-. .i -, r 'S '* '- *i ; '-'
= f I:- :"-r--!"--"T^T'T^ T" *i--^ -" -- "[" I-


4'I


I4ff


-I

I_:-
V

1*


itFtI


r 7--I


^^::?-
i I







1-+-




--






-.


"- T -T'- 7 t + .-..

_- J- ,--i-'4 -i+-_
!- _^^tr--li


A : tt- .c'-----i-- i
__.- -: -._ t


.'.~~ ~ '( ;"
7-4:t; :-,
-tiT .--
+tt : ; ii L
4


-_--C; 4-i


- p 7 '-

Lqzb


F-f
ii -t!~"3


aTE~LWVIIt __Ta
2J^^fiufl^^-P7R" 7- +


I


I J I* L I I I-- 1 : 1 1
s 9 .^ I I A L~li:


, ~l._...


~7 r~r


": I-, "1 T1 = s .. -T


~FT~C~F'
i
_ILL~L~I~L~ !~:


..........-- -- "'. ...- .- i-r
,L. .. -. r' .- _. -


p -'.. "


- .--. I -.". p --


I


11 I I N L i __


I


. '`' I-U L-. -. .. ... .. I L ~_I I I I ~ Y


i


i


-T----- 1- P --I ,- 1+ i,- I --I I I4 -


Fig. 10f


I .- 1 I. ]
-4-L
r-4 :4_1-r-


;ff


r i
Y; `


7


i;

~I ~


j L, ,.-

.i -








NACA ARR No. L5H09


Fig. lOg








NACA ARR No. L5H09


-- "orce-lest daQ(
- -- Ca/cu/atecd from surveys


0Z

0

02


/l2


-I


.?

(


8 -02
'--:
r)




. 0

0

' -02

-[i

N


CL04
1.04


.83





.80





.40






.283


-/0 -5 0
Ang/e of yavv ,


5
?~, deg


/0 /5


(a) f 0.
Comparison of ,icremen! of Cn due to
verficol fQi/ as determined from the force-
test datf and from the qir-flow surveys.
Propeller removed.


92---- EE EE i:: ::
02



O -8.4

orr
3. 4



NATIONAL ADVISORY --
-- .COMMITTEE FOR AERONAUTICS-
-.--- -::i: L=:::;- :: i:::: l!
:O ::1::::::::::::I I I I I I
2piicm mm mziI


-15


Fiqure II. -


1 11 1111 E 4443---LIuII II ll Il l'


,


Fig. lla








NACA ARR No. L5H09


SForce -es dato
---- Co/culated from surveys


1 1 I 1 1


---


ItiI


II


a

1/8


H M M M -M I I I I rl





928


1


IJ-I r I liiln rr ~I 11I I I t1IIIr


411 11 111 ii11]IiI


S I. I i I.. I t


SII I I I I I


NATIONAL ADVISORY
-COMMITTEE FOR AERONAUTICS-j-


I. J


-'5 -/0


-5 0 5 /0 /5


Angle of yaw 7, /dey


(b) 1f, Sco0.
Figure // .- Concluded.


.0
0 1
1L,


/56










1.II


.C


0


-.0I


F


k..9 1/d


Fig. lib


--


-:<


I
I


,


I I I I







NACA ARR No. LF.09


- force-fest
.---- Ca/culated


dcto
from surveys


_I 4L.-.---


9.


(dog) C.
3 .804





Z .83


C'.)


















14







43
C'
C'
N.
C'.






I-

Q)
I-


IIItlil


3.4 .40


m zI I I Ill


/0

NATIONAL ADVISORY
- COITTEE FO AERONAUTICS -


-,'5 -0 -5


0 5 /0


Ang/e of yaw, 74, deg


(a) SL,0.
Comparison of increment of Cy
fail as determined from the
doao and from the oir-f/ow
Prope//er removed.


due to vertical
force est
surveys.


S8.9 .80


Figure /2.-


n1lrn111717rllIr1 111 I-1 I i I I I I I


Fig. 12a








NACA ARR No. L5H09


- Force-tesf data
--- -Ca/cu/ated from surveys


oc
(deg)
/1.8


-/0 -5 0
Ang/e of yaw,


.6
1.56


(J








]
I0

!





U
U
0
I



'N

K
0







U


NATIONAL ADVISORY
COMMITTEE FR AERONAUTICS-
5 /O /5
76, deg


(b) fr, SO".
Fiqure 12.- Concluded.


5.8 1./








4.9 X04


-1 L
-"'5


'I I I lii ii I ii 111111


Fig. 12b







NACA ARR No. L5HO9


I I I L I I


.111.11111111~~~+


- iih Flaps def/ected 50Z\
: ,/ir ii: III 111n 1


rIIzI


Il l


I I I I I I I I I I I I


~I


cC
~Jeg)
4.3
1.8











(deg)


/. O
,3.4
-8.9
-9.2
-112.3


t it H Itlt H Hi if ill H i
- _Z _l !! I!- I !- .v,/ i/


F/ops refrocfed
NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS
S1 i i 1 l l, ri


5-


-/0 -5 0 5 /0 /5


Ang/e of yoa, j7 deg

Figure /3.- Variatfon of air-flow factor with
anqle of yaw. Propeller removed.


-/5
-7


CL
1/.04
/.J6
1.,F6











CL
0.23
.40
.80
.83
1.04


II i llI


q7aaa~a


Ntoo
,GjZ%

_5s


Fig. 13








NACA ARR No. L5HO9


1 I I:I :


oc
(de g)
//.o
,,3.4
Y.-R


! I I -3


i-jj


III J ill 1 1 1 1 1


0c, Oeg
/.0
------ .4
--- -- 8.9
----9.2
-----f2.3


Power Lift coeff//ent, C,
Rafed0( =.056) 0.24
Raed(T= O.//) .43
Rated(- = 0.30) .96
T = O/ .83
Rated( = 0.5/) I.39


i iI i i i i i i i i I


SI I I I I I I I I I iI I I I 11 .rr i.~


!l lli lli l lil ii I


NATIONAL ADVISORY -
COMMITTEE FOI AERONAUTICS -
l I I liiI


-/0 -5 0 5 /0
Ang/e of yow ?", deg
(a) 4f, 0


Figure /4. Effects of propeller operation on the
d', ecftiona/- stabi//ty charocter/if/cs of the XF6F-4
airplane .


0


.02


.04/
-/5


oc
(deg)
-/.
-3.4
8.9
'9.2
~/2.3


I l l .I ..Irtl I I


2 J=F


ztl 1U 17T:-- T


i-


Fig. 14a








NACA ARR No. L5H09


O



o






oC
I
a








( -
o
I,
^1













li <^
0
90
1.1b


1^~

is'


H Li1.


Cc
(deg)
-.49
=-//.


or, deq Power Lh'f coeffc/irenf,C,
4.9 Roated (TC =O./) /.39
5 0.65 Raed (7-=0.33) /.37
S //.8 T = 0.0/ /1.5


NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS


.04r


cc
(deg)
--7/.8
--5.8
-4.9


SI I I I I I I I I I I I I I I I I I I I I I I I I I


-0 -5 0 5 0/ /5
Ang/e of yaw, 71, deg

(b) djf, 50.


Fqure 14. Conc/leed.


-/;41


I I iIl 111


. -f . . ..l i I


11111111111111111111111111t111


. .


=:4 4= I-I
L4-T I


-.02


53


Fig. 14b







NACA ARR No. L5HO9


- force -test data
- -- Caculoted (Prope//er a/one)


oc

1/23
1 9)}
\IPl.1


C,
/139


0.5/


I L 1 It r
II J! 11 -1 .E


i 9.2 .83 .0/
AI I HI4--14


i. i I ii.i i.a.
IIL Ht I
i l I II tl '


.96 .30


I I3.4 .43 .11
; I I-T


!: : :.....: ..: : :: :.


.24 .05


NATIONAL ADVISORY
COMMITTEE Fm AERONWATICS


-5 -/0 -5 0
Angle of yaw,


5 /0 /5


fa) df,00.
Figure i.- Exper/inenta/ and ca/cu/ated effects of
prope//er operat/bn on the var'/f7bn of Cy w/th
31 for the ai/-plane with vertical/ tai/ removed.


! 1 1! 1! I I 1 1 1 1 1 1 1 1 1 11 11


MI IJ1-: +-FHT-


~


i


tttffKKKKtftt~


'''''''''"''''"''"" ~"'"''


Fig. 15a







NACA ARR No. L5H09


--- rce -feas doto
---- Calculated (Prope//er alone)


ac
(deg) C
:i:t:::::::l:ti:l :t:://.8 158


,-111 1


-I./ --I I I 1 I I I -


137


II1*T~1


S- -i I I I I I 1 1 I I I H1 I I lI I I



0


NATIONAL ADVISORY -
COMMITTEE FO I AERONAUTICS

-/1 -/O -5 0 5 /0 /5
Any/e of yaw, 3, deg

(b) 6, 50
Figure /s Concluded.


QO


/.39 .5/


I I IJL


PR;3)~


I- l l l1l111F11l 1 1--


Fig. 15b







NACA ARR No. L5H09


Force -tesf doat
---- Calculoaed (Prope/ler a/one)


oc
deg) C-
i I i i l I I I I l I II I i I/ /.39
!!: :: : !!: :: : :: : ::! /J5-_._-tiL


0.51


.83 .01


a' -


.96 30


.43 ./


. ... i i I r


.24 .05


-i -5 0
Ang/e of yaw,

(a) cf,0.


5 /0 /5


FIgure 6. Expe.r/ienta/ and ca/cu/aoed effects of
pr loe//er opeerca/n on the variation of C,
with 3 for tne airplane w//h vert/ca/ fa/1 removed.


-I


5


;--.;--


'I '4


1 I i 1 1 1 1 1 1 1 1 1 1 11 1 I1 1


1 I I I 1 11 1! !


II


I iI


14- -1 L4


I_ 1 l 1 1 1l l i 1 1 1 1 1 1 I lI I I


!r


7=Z=F-


NATIONAL ADVISORY -
COMNITTEE FOA AERONAUTICS
i-- c1 -l -l I


-.02
.O2


Fig. 16a








NACA ARR No. L5H09


SForce -fest dofo
---- COacu/loed (Prope//lef a/oe)


(deg)
: //.8S


CL T

1.58 00/


LI I I i i I i II


Q.


(:


x







A,
. -J





V


0




91~
0
\a
j















V
V


rT


II II I Ii111111111 II II II I


- NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS


-15 -0O -5 0 5
An/le of yaw, V, 'de_

(b) ,o 50 .

Fgure /6. Concuced.


68 /.3J7 33


/.39 .51


/0 /5


0

-02


II I T I III ji l li 171171111


14


. .


4.,Y


Fig. 16b








NACA ARR No. L5H09 Fig. 17a






L- -- --- "i .. ....: -"- i ... ...
S;F: r:- ; '-1: -
.... I- i y. r r. ,. /





1y-
F 1 -... ..- i -'' I.. I'-
.... ., t, .- t- ... -:I





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-- L i



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I I
t' -- t




ri

r
r--- --- .. -',- .. .. .... ... -'-- +b -- -
I- L .... I-- _t---:--- '- ..-I-- -, '



/ I- il

,- ,- _
.....~ ~ ~~- k >- -



._. -. ----- '_.





I-- __i I I -



/ --











NACA ARR No. L5H09


a~rv -r~iaALL I44iric.


Fig. 17b





-, t -- I-' l"---l '- ---, I -


!-I
+-4T r




,+- J r, .
-- -










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.........

















-44-
- -- r n t- -- "- -i--
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----- --,----- --' ---











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----- =- I--























LL.
-~ -- -
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-1
-1 '-___ __"-



-r r -- ---- ---
I.'P II 1 1









...... -- __ i,-- _-^ -_ --- -'- -- --=- '__^-i -
- _- -_--,- ..:-- _.' ..... ,. ,-K .








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.-.I-







z _. --- __ :_ .__ -:-,:- ---


4.-- --


.1


. r- .- l--
- .... .. -- _
l~-
. ... L .








NACA ARR No. L5H09 Pig. 17c





-- -- -" -

*- .. .. -- 1-- 1 -!
. [* .. ..- i .- t ..


+, -- -;. -. -; : t -..- :- ',-+ : -i .






-I




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j 4- .. ~ ~i l ..



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^ ^ '.f:-~:---~-;-.- *--^^^ ."-- ---- l "\- i '- **- ^-i}J
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.. .... ----- : '-- ---- :- I -_ i _, :
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'I '-' l -_
.... ~ ~ ~ -- +. -.--- ,-__ .r ...:-.-_--_-- ... ._- _-. ..._'. ...__






--- -- --, .-. .-... 1 ...




? ------ -
-l--- -- -
'+' -- -" .. t .. I .. .









Fig. 17d


NACA ARR No. L5H09


S .. _
S.. ---:I-

-7 .- -

r- I I "I


---t-








I --
. t -




. ... I_---

.- .-. 1 ..
S-. *-- --- .

L-- -J .... : _: -- ; -,-:T
1- "-ii



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.. _-^ .- t
I-... -

41 4




-!- -:- :-I-I -:- ;
-!L .i- I:..-7:, -At44


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- -Ir--I--


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-V- -





- -- -n-



--
--~ '

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--
--_ L- --
-L
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-i- -i-r-


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.:,~~~ ~ .I t -; ,

-- -:-:.- :- :








----t :t"^S-Yi !

---i --
:--- -I--- i-:


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;_ --: -:- ..L ~ :r* > ~ ', ~ ~ i --L : -----7-_-, ------ : 7- :: -I t- -- :- ----

. 1 .__ __ ___ __-_-----i-i----.



L-N

- f : ___i



I- r I -.
_ -1-_ -. P- _.
I-i ?4- 4:-.$rqT


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!:- _.r- -- .--
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-- ----


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ILI
i-9















--I
----~--I -


i
1


-. ^-.- -_
1's 73t-- T
.. T E ..









NACA ARR No. L5H09


Fig. 17e


h -- ----- j! -- -*- i-

S '-i
S -. I
":A- D --

F I K.
L- : -_ .. .

-- t -- --
... --_
i4-
I :--t:


C- -~



7:




-'---: _.1 f I. _
4~-~ pTf~,


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1-^-~-


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. _
I-.
i ...
- .... .. -- i."
- --I .. ;--i -.
hS- -+ ._ .- + _

t: 1 : : --+.

^--'-. _,,_ _--
-~: GL^ .f. C


-r-T


-_r_-I- -
S -- ? --- -- .. -- _...

--- --- ..."-H .. .. F-- $- 1 -t t-' -. i- 7- -








-II
S f 4Q


ZF








44
.. ..... ......... ...... -- -' i- ,r'- ^ -- F--











_._.. .-_. | ,_ .. ... .... .. r -_'_- ---




j ZI--,& I .-
: -- '
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r --- _- .. r .. ... .. ... .
I *--- l- 1.,---- .. ... ._













.. ri< ... o ,- "- '
-- .. -- | + '


.. ... .,--- -- ,. . ..









NACA ARR No. L5H09 Fig. 17f









S- 4- -- -

". .- .. : .
-I ,r -j I-
......... :..-TT ; ___ ."-I i- [l .-- r l b
--- ------ ----.. .... -









i -


















.- -- ,,. : -- L .. T

-. ~ ~ ...'.. .... .. .. ... .. .. .
-- / -
7- e-' --- :---l-------
4-1i -K -






L4-








NACA ARR No. L5H09


K -


-.. .. -

L :--:: i.


I -


i--I


I" -
... r ..... "_" -..-. ^ i i L .._..



-t- a T, 1 h-- --, --
I -" I
..... i '..

.... .. '-. ; "...,-I--C t-'. i.
--. 4-. -. .. -- -- -- -- -- ----I-- -.-i
_-- -." -,i -- I .
&

'4-- f--


.. .. ... ; ... ,> ... .I j.F


-* I -- L __ t ; p- t -,: -- 4 -. _
t 4t I~ ciii
-t -" .. I ... -.. ... ..
.: .-_ ; .... ; ,- ___ ,:' .__ .-

'. .I. .. _
"---- "- --- --,- I- -- -; : -_

'~" ~ '. '^ ^ ^ .. !- iL ..., i --I -.. ......
^ ... .. .- .- ; I, '
-'- _-L -- .,-:---,--; ... --, .-- --H-.- .- .----- C-- r--- i
._ ... _.. _t ._ .; .' ,__ .. ;... I .. : t-. c .- -

S--- .----- -: L -- I--

-I : ---" .- :% ... *- i--" -'-. :, I --, <- '- :---
-.... .. ; d .. ,-- .... .... t -, ".. --.
i- '- .. :_ ..L _.- ._ _: i _- .- -' : : : !--^ 1
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L--''- I ,,' .
r ', I .
-- .. .. ... -

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.- .T .. .. -
4- +
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,. ,, L r'
SI -

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-;-- --,I
-f I


' -I ,< i C i-
I C, ... r--
Kt t?
rn7y; arC~-~TE :u~'


Fig. 17g


j


'''










Fig. 17h


NACA ARR No. L5H09


1- -----F
-.. .. i : .


.- ..- .. lt r .

,- -. _r -- ..

-- y ^ --- ,-g -7 -^ _-^ _^ ~- *r-;-- : : _;: :*;

S_ 7 -T -S ~,
S--


_I r U vv- ........ .. 1 l -- '- ItIr.. .


illi I. ... .. ll

_ I_...I .
:- --- : -r-

- I-., -II2- i^-.'-i



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_- _-_


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-C.
I-' '


" -4--r--'

- --7
-;-- I- F^--




.~?-- $ ..... -l.. u-_
LL_
_--- h l I- --.




:^ L- = ^ q--'q.L "
:-*--- -: ,- 7 f /


Olll'] fi l -- L -" --
l I I I

I-
I-i~


' '-- r 'l, 1








t -







- -- -
l II il

rnjtt^


r-r-
.-T -4


S-- I- -l -

-.^- --"
H-- :







-l-- --l

44-t


- .----- -- I1- J' I / I' H-- :' rr
---- 7
-1


S t- I-- l--
'-V ~ -:-,4-" -_-t -I --- Ht-~ -Fy I
r''v----i


: 1 1-- _
zrsi a *Cr'v'r-1 -' -
-- .-
Zar 5a" -" -.
T.


- ------t-i- ------ --

I -l-- -'''-
--- r r -.
-- f I


fr-~II


---r---I
-4-




---I-

:rztr 4
-.


7 -


I 1 .I 1 ''I ,


L


-, -----
-"mp~ --

'7
- + ~-Lt --
cL!' '
I='F~--
h
--C
-t--







NACA ARR No. L5H09


(oeg)
12.3
8.9
9.2
3.4
/.0


Lift coeff/cienf, C6
0.24
.43
.96
.83
/.39


Power
Rated(c'=0.05)
Rated(E=O.//)
Rated( c=0.30)
Tc =0.0/
Rated (Tc=0. 51)


Figure /5.- Effect of the prope//er s//p;srean on the
contribution of the vert/ic/ fa// to the aop/o/ane
d/rect/bIo/ stabd/ity c/aracter/"f/tcs.


oc, dg

----- -13.40
3.4
---8.9
---- 9.2
-----/2.3


K
^1.


(3



















K
CI


.02

O r
0-


.02
-I


NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS



9.2
-1.0
8.9
/2.3


-/0 -b /0 /$
Ang/e of yaw, ?, deg

(a) .


I L. L I I

i i Tt


Fig. 18a








NACA ARR No. L5H09


oc
(deg)


ii i:: II I


~




s
\c:
P,
a


9P
,


C,)
0)


K-5.8


Lift coeff/cfict, C


4.9


.---- .8 0.65Rated (c=0.33)


-- /11.8


-/0 -5


Ang/e of yaw, 7 dey

(b) 6f, 550

F-igure /<. Conc/uded.


Rtoed (To=0.5/


T =O.0/


/.J9
1.37
1.58


o





1NZ
i


-' '


C,)


NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS


0 ---- -


no z-


c
(deg)
//.8

--4.9


-/5


0 5 /0 /5


I -1 -


ii i


Ifc


I 1 1 111 I I ll i l I


x, deg POwer


-f


Fig. 18b








NACA ARR No. L5H09


/-orce-fest fdofo
- --- Ca/cuated from


surveys


oC
(deg) CL Tc
/Z.3 /.39 0.51


4- I


19.2 .83 .0/


1













I,
t3


i i I I I I I I .9 .96 .30
A l J l l 11.. . .


I 1 i i 1 Ti i 1i I t t --L 113.4 .43 ./


NATIONAL ADVISORY
COMMITTEE FOI AERONAUTICS


1 I


. 0 .24 .05


-:0 -5 0 5
APg'e of yaw, t, deg


ao;, 0o.

Fj7 ue 19. Comnoarison of increment of Cn due to
verhicoi' foil/ s determined from the force-
-est dafo and from the air-flow sur ys.
Propeller operoainy.


0

-._0

-.0
.02


-02


oz


-4--:t I


-.02


-* I I I


"""; ;4 -,,,,,. [ I I


I


-5


0,


.1ig. 19a







NACA ARR No. L5H09


--- Force-test data
---- Ca/culated from surveys


I irmt i44ri- I I


Ii: III11II*II*


/H.9


cx
ieg) C ,
1/. /.58 0.0/








58 /37 .33


-/5 -/O0 -5


0 5 0 /5


Ang/e of yaw, d' eg

(b) 65, so0

F/7uare 19. CorC/Lded .


.04


.02


~r.-1...


I I I 1 -


I I I M


tttttttttt+ttttttttt~trt+tt~fl


I I 1 1 t


-4L IIII I I I I I I I I Ii-i I I I II


rllll 1 M M I 1 !I'll M I Il ll~ffl'lm


-1 1 1 1 1 1 11 1 1 !.: =


Ilil I rrrn


~f~''' fFttt~KKHtttttttttl


-Of--j


NATIONAL ADVISOY -
COMMITTEE FI AERONAUTICS
1,I i .. .. .


-.02


.02


Fig. 19b







NACA ARR No. L5H09


-- Force test dafa
--- -Calcu/ated from surveys


cx
(deg) CQ TC
/2.3 1.39 0.5






9.2 .83 .0/


N









0,
U
0
QJ


'if
0,

U


1II lIJI I*lllI ll Ij7111111 .1III


3.4 .43 .//


....-. H....2/. /0 .24 .0


NATIONAL ADVISORY
COMMITTEE FO AERONAUTICS
* a liii


-,5


-/0 -5 0 5
Ang/e of yaw, 1, deg


/0 /5


(a) 7,, o.

Fg&'re ZO.-Compaor/on of screen ofCy due to
verticol fail as determined from the force-
tesf dafa and from /he air-f/ow surveys.
Propeller operating.


uzmzII


:8.9: .96 0J


ttli [iL T


-I


I I I r II I I I II I I I Ilil I II I I I II I ~rl


EEI


T1


1 1 1 1 11 1 1 1 1 1 a :1


1I1I 1 II l ll l I I I. I


tmn


tttt-ttttttttttt't~:~


a:


-111111 11111I 1 11 1 1 =III I


I+Kf~


I I I lI-4-- +- t---T III III=


II =l I4 .


' ~ ~ ' ~ '


s l


Fig. 20a







NACA ARR No. L5H09


SForce -fest doaf
---- Co/cu/ote'd from surveys


i I I 1LLLJL24t I 4 -- M471?


ID


J. I I


cx
deg) C2 7T
/.8 /158 0.0/







.8 1.37 .33


NATIONAL ADVISORY -
COMMITTEE FOI AMEONAUTICS
- I 1 I I 1 1 1i


-/0 -5 0 5 /0 /5
Ang/e of yaw, 1', deg

(b) 6, ,50


Figure ZO. Conc/uded.


-/


-/5


11111111111111MIIIIIIIIIIIIIII


i I ill I I I I I I i I I I I Irrrr-~rll~


3T


Prt$ttFfKFfn


4.9 139 .5/


Fig. 20b







NACA ARR No. L5H09


doc oeg
1.0
----- 3,4
--- 8.9
-- 9.2
- -- -- I.3


Power
Rated (c=0.05)
Rdated (7E 0. //)
RofCed (7Z 0.30)
at = 0.0/
Roted ( c O0.5"l)


N4444 J44 m~m-rFTTTVW I VW


Liftcoefa/ent, C
0.24
.43
.96
.83
/.39


(deg)
-/.J
'-8.9


OO
















-/O




NATIONAL ADVISORY
COnnITTEE FR AERONAUTICS
-0 -5 O 5 /O -9








Ar e ofyow, #-, 'eg

(a) d," 0_

F/oLJre 2'.- Val-r/7/bn of ar- f/ow factor w;ith
arge "' yaw Prope//er ope/ra/-.
-5 -0 -0 5 /0 /












ang!e or yaw Propeller operatl/n .


K

"U


NJ


TTE


I I I I I I I I rl -r 117rlKI Irlilrnl


Fig. 21a








NACA ARR No. L5H09


Po wer
Rated ( =0.5/)
65 Roed(T = 0.33)
=Tc 0.0/


11111111111111


'L'Mt ll 1 litilil 11


1- ll I I 1 I I. I .


I I I1 l
,


III


zut-m-


SmiII


I I 1


II 1 1111 1 I '1 I II I I I
..i .. .- -


0c, o'eg
4.9
----- .8 c
----- 8
---11.8


mI I


S i I I I Ill I V


/ImI


SIItttr 1 I


tr


. . .


III rII,


..lII .l III


zIzrzmzrztttn I


IIII lzlzi 111I


- tI I NATIONAL ADVISORY
COMMITTEE FO AEONAUTICS
sI I t I I I I I


Ang/e of yaw, 7, deg

(b) f, 50.

Figure Z/. Conc/udecd.


L/ft coeff/cen, t, C
/-3L

/.37
/.56








-5.8
(deg)


fi i 1 f II I I I I i


5





S-5


r/lY ~II i ~l I I llI I


-2C


-/5


-/0 -5


0 5 /0 /5


iii I 1 17!1 1 I I I r


....... V M 111
1.11A 1 14 1 1 11
~ -1 / 7 i ,


I I I I I I I II i rT


.It


. .1 1 1.. 1 I l l .Il l I


,,,,,, ,,~


. I I I I I I I.I.. I I I
I/ I/I I I I-I --" -


E II I I I I


. .1


1:1I11 1~171 1 i I i 11 [711171 1 lrl I T I


I I 1 1 lt


I I. l I7


Fig. 21b








NACA ARR No. L5H09


-20 -/O


O /0 0O 30 V
SI |I I I (deg)


K)


u


I


N
K

K.

N


A- -


-/4.6
-99
O
/0,0
/47


NATIONAL ADVISORY
I COMMITTEE FOfR AERONAUTICS
30 -20 -/O 0 /0 20 34
Rudder def/ec ion r g ,

c0)Gidlng con.d/8bn; Tc0 00/; 0t 9.2
Ce, 0.83; f 0 .


Figure 22. Variho/bn or C, and Cy with dr for
several ang/es of yaw. Propeller opera t/ng.


-30
.2 r-


Angle of yaw, K, deg


-k



02
^' 0-


.02 -
N -


.04
S-


Fig. 22a






NACA ARR No. L5H09


-30 -20 -/0
.2 tII--,---


0r



N.


(dleg)
--/4.6
--9 9
--S.I


(b) Cl/mb/ng conditIbn ; rafed power (Tc=O.,30);oc,8 ~9
C, ,0.96 ; d6,o.0

F/7ure 22.- Cont/n ued.


0 /0 20 30 v
(deg)
/4 7
/ 00
o00

-146


- -

' -04-



, -30


-20 -10 0 /0 20
Ru dder def/ecl/on -r deg


Fig. 22b








NACA ARR No. L5H09


-20 -/0


0 /0 20 30


--o-- -/4.6
----- -9.9
o-- -,F/
-- 5.0
/0.0
------ 14.7


-20 -/0 0 /0 20 30


/rudder def/ecft/oln I ,J de7

(c) C//~inbg conoii/on ; raTed power (7 =0.5/);
X, /2.3 CL, i.39; 6f,0.


Figure 22 Cont/inued.


(deg)


-30




./-


0-




72/ -

:2 -


K)


N.

C)
0



U

'N
VJ,
Ns


-30


Fig. 22c








NACA ARR No. L5H09


-30 -20 -/0


S /0 20 (d0g)

T? -+-- /147


/ngle of yaw, w de
--0---- -/49
------ 14.6
-- 9.9
-- -,5..1
0
,5:.0
----- /00









/-/7


4-9.9

NATIONAL ADVISORY --- -
COMMITTEE FOR AERONAUTICS O


(d) Wave-off conoa/f/on; raTed power (T=0.5/);
c, 4.9 0 C,, .3S ;ff,50.

Figure 2 .- Cont/nued.


0


S04


:06
-.


-20 -/O 0 /O 20
Ru dder de//e o/on r- deg


.0


Fig. 22d








NACA ARR No. L5H09


-20 -/O


0 /O 20 30

1 I I I I I I (deg)


r,.





V
K




NJ
',






N
K



...

u,
I,t
*I












'3
*^
cs

Fs


-30
.2 I


-30 -20 -/0 0 /0 20 30

Rudder def/ec /on -r dey

(e) Landing condition ; T, 0. /;11a, C,,/.58;

F.ure 22. -Conc/ue .
/fl>re 2.2. -Concluded.


--o-- -/4.6
-e "9.9
------ .I

-I/00
---- 147


Fig. 22e








NACA ARR No. L5H09


'IZ
I.2





0




















.o0
"2





'.d


i .02






^ 102
I^

\z-O


x


-30


Angle of yaw, y, deg
---- -4.6
-- 9.9
-- 0
0
---- /0.0
14.7


1 I I NATIONAL ADVISORY --/. 7
1 COMMITTEE Fil AER ONAUTICS
:04 '
-30 -20 -/O 0 /0 20 30
R dder def/ec ion, Fr de!

(a)it,9.20; C,,0.8 ; d7f, 0
Figure 23. Var/ tfon of C, and C, with drr for

several/ ang/es of yaw Propeller removed.


eig. 23a









NACA ARR No. L5H09


-30 -20


Fig. 23b


-/O 0 /0 20 30


tc)
N,

QI









S.




N~


-20 -/O 0 /0 20
Rudder def/ecfoon S dey


b)J a,/3.o0 g ,/0; 6,-, 0


Fgiure 23. -Contlu/ed .


----- -/46

------- -/J?
- 0
------ SO
S I00
--- 147


.02


F;02
1 -.02




- .04
x^S.


v
(dceg)
,-/46
--99
- -S.I
--0

/ao
- /47


-30


NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS
!MITT.


rr,


^








NACA ARR No. L5H09


-20 -/0


0 /O 20 30


---- -14.6
----- -9.9

6 5:0
--a- -,.I
----- 0

--o- 147


NATIONAL .ADVISORY
aI ___ _COMMITTEE FOI AERONAUTICS
-30 -20 -/O 0 /0 20 30
RuAdder de//ec/onr/, deg
(C) a, 5.:6; q, /.09; O50


v
(deg)
--/46
--9.9
--s.I
-0
- 00
- 4.7
-/l4.7


Figure 2.- Continuedd.


-30







I -


(de
(deg)


Fig. 23c








NACA ARR No. L5H09


-20 -/O


0 /0 20 30


Ang/e of yaw, F, deq


---0--


--------


-/4.6
-9.9

0
.0
/0.0
/4.7


K .04


s.0


0.0


~- -
-.02



S-30


NATIONAL ADVISORY -
COMMITTEE F01t AERONAUTICS
-20 -/O 0 /0 20 30
Rudder def/eo/on r-, dey

(d) ac,1.8 CL, /56; cL,50


Figure 23.- Concluded.


-30
.2 r


Fig. 23d







NACA ARR No. L5H09



S/





o ,,9.2 ; C0O.83 ; propel/er
----to,9.2;C,0o.8; 7c,0.0/


20 11 I I I I I I


Fig. 24a


removed


1;

it
Xz,


-I0

-20


I I I I I I I '* 1 1 1 1" I. I
-1/ -10 -5 0 5 /0 /5
Ang/e of yaw 7", deg

la)G6/1'gq cond/ft/bn ; df, 0

Figure 24.- Directlono/ trim characteristics
of the XF6F-4 a/rp/ane .








NACA ARR No. L5H09


./ I-


0
It1


r l1


a, 9.2, C, 0.83 ; proped/er remo ved
-- -,6(.o9 q, ,0.96; rafed power(^=0.30)




'T, ~ ~ FF __-- _


0


)1 1 1 1 1 1 1 1 1 1 1 1 1 1 -l m


rI


It


|,-1:-+ 1 1 1 1 1 1 1 1 1 1 1 1


111 111111


I' -- ii i hi i i Li.--- .


Ang/e of yaw,


I III I I
NATIONAL ADVISORY
COMMITTEE FOa AERONAUTICS


o 5 /0 /5


71r, deg


(bl G/imbing condI/'bn ; 0O.

FIgure 4 .- Continued.


-.0


(


II


-/5


-40 -5


Lm


I I .. .. .


I I


,--


Fig. 24b







NACA ARR No. L5H09


Fig. 24c


.I
0
It


I r II Iii* : 1 11


I I 1 1 11


propeller removed
rated power (=0. 5/)


0
II


1 111 l-i --ii i


II I I I
NATIONAL ADVISORY
COMMITTEE FOm AERONAUTICS


S O -5 0 5 /0 /5
Ang/e ofyow, -, deg

(c) c/nmb/ng condi/'bn ; dc 0


Figure 24. Cont/01?ed.


-.00
-w-OO/-^^ ^


-L -- L- L L14-r

=I TT T7 T


'0
-- I-I-- I- --L 1 1111
I I _

^-EE^EEE^EEEEEEE L IEEE


SOc,/3.0 ; C,,//08;
- C,2 .30; C,,/.39;


dl k








NACA ARR No. L5H09


It
NU


0
II
-c


.4,












a- -56 ; C4, 109; prope//er removed
--- 0,, 4.96;C,,/39; rated power (=0o.5/)

20


/O





-0


-^:::::::::oE:z:~^


-3011 i i 1 1111111 1 i I i i i ll[


Iz~l
(j ~


-001


I I


I I~__I II I NATIONAL ADVISORY
COMMITTEE FOO AERONAUTICS


15 -/0 -5 0 5 /O /5
Ang/e of yaw, 7, dey

(d) Wave -off condi/7on ; df-, 50.


Figure 24.- Gont/hued.


1


Fig. 24d








NACA ARR No. L5H09


./
O
0
II


ao


J.


/O

-0



-10


-201 1 zz


|11 .


IJI4LWH HII LLIU=i M
I I -1 f- I I I I I 1 1 1 1 1 1 1 1 I I I 4 -


O


NATIONAL ADVISORY
COMMITTEE foR AMRONAUTICS-
-/ /0 5I0 5 / / II


-15 -/10 -5 0 5 /O /5
Ang/e of yaw r, deg

(e) Landalg Condition ; cdf, 50

F/gure 24. Concluded.


C ,/ J6; prope/ler remo ved
C,,/.583 O.O/


_ _-1- -I- _L


---- ------- -


~1 111 1 111 1


~Li I i I i I I i I II I i I II1II Il I I I i I I I T1


I


Fig. 24e








UNIVERSITY OF FLORIDA

3 1262 08104 965 1


UNIVERSITY OF FLORIDA :
),,DIJMENTS DEPARTMENT
1 'MARSTON SCIENCE LIBRARY
O. E DX 117011
C.J; JESVILLE, FL 32611-7011 USA























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