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*7-RESEARCH
INVESTIGATION
.
l%FOOT
‘J.
RANGE
OF A THIN WING OF ASPECT
PRESSURE
STABILITY
MEMORANDUM
WIND TUNNEL.
AN3 CONTROL
OF A SEMISPAN
By Ben H. Johnson,
Ames
V - STATIC
THROUGHOUT’
MODEL
RATIO
4 llff TEE AMES
LONGITUDINAL
THE SUBSONIC SPEED
OF A SUPERSONIC
Jr. , and Francis
Aeronautical
Moffett Field,
AIRPLANE
W. Rollins
Laboratory
Caw .
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4
NATIONAL ADVISORY COMMITTEE
FOR AERONAUTICS
WASHINGTON
December
8, 1949
TECH LIBRARY KAFB, NM
NAC!ARM A9101
NATIONAL ADVISORY COMMITI’EXFOR AERomKJTms
RESEARCH MEMORANDUM
INVESTIGATION OF A ‘IH13Y
~
OF ASPECT RATIo4m
Tfm AMEs
12-OOT PRE3SUlZEWII?DTUNNEL. V - STATIC LONGITUINXKG
STABILITY AND CONTROL TEROUXOUT THE SUESONIC SPEED
RANGE OF A SEMISPAN MODEL OF A SUPEFSONK! AIRX’LANE
By Ben E. Johnson, Jr., and Francis W. Rollins
STJNMARY
A
#-.
d
Wind-tunnel tests have been made of a semispan model of a hype
thetioal supersonic airplane to determine the static longitudinal–
stabilityand -oontrol Oharaoteristics of the airplane throughout the
range of subsonic Mach nmibers up to 0.95. The semispan model had a
long slender fuselage and a wing “and horizontal tail of asyect ratio 4
and taper ratio 0.5. The midohord lines of the wing and of the hori–
zontal tail were normal to the @ane of symliletry.The profile of the
wing and of the tail was a sharp-edged, faired, symmetririaldouble
wedge with a thickness-ohord ratio of O.042. Tests were made with the
horizontal tail nmunted in the extended wing-chord plane and alternately
6$).
6 percent of the wing mean aero-c
chord above the extended wing–
chord planet At a constant Reynolds nuuiberof 2,000,000 measurements
were made with various stabilizer angles of the lift, drag, and pitching
?mxnentof the model at &ch numbers from 0.20 to 0.95. With the wing
*
flaps deflected for maximum lift, similar measurements were made at a
Mach nuniberof 0.20 with Reynolds nunbers up to 10,000,000. Measurements
were made of the dynamic pressure at the two locations of the horizontal
tail and of the character and location of the wing wake for the range of
l+kchnumbers and Reynolds numibersnoted above.
At zero lift~ the Mach nurib
er for drag divergence, defined as the
Mach number at which the slope of the drag coefficient with respect to
Mach number equals 0.10, waa about 0.92 for either looation of the horizontal tail. The angle of attack for a constant lift coefficient
decreased slightly with increasing Mach number but no marked or abrupt
compressibility effects were evident at lift,coefficients less than 0.6.
The contribution of the horizontal tail to the static longitudinal
stability at low lift coefficients deoreased with increasing Mach nuder,
/’ mH?”
2
RACA RMA9101
primarily due to an increase with Mach number of the rate of change of
effective downwash angle with angle of attack. For the model with the
horizontal tail in the extended wi&-chord plane, this decrease in the
contribution of the horizontal tail to the static longitudinal stability
was aggravated by the reduotion with increasing Mach nuniberin the
-C-pressure
ratio at the tiil. With the horizontal tail mounted in
the extended wing-chord plane, static longitudinal stability existed
about the quarter point of the wing mean aerodynamic chord at all lift
coefficients for &ch nunbers less than 0.87. At ~oh numbers between
0.87 and 0.95, the model was neutrally stable or unstable at lift
coefficientsless than 0.30. With the horizontal tail mounted above the
extended wing-chord plane, the results indicated static longitudinal
s,tibilityat all lift coefficients for all M@h nunibersfor whioh &ta
were obtained, For both positions of the tail, either an all+vable
stabilizer or a constan~hord elevator provided sufficient longitudinal
control to bahnce the airplane at all test Maoh nunibers.
t
INTRODUCTION
.
AS a part of a geneml program to determine the subsonio charactez=
ii3tics of wing plan forms suitable for flight at supersonic speeds, a
series of tests of a thin sha~ged
wing having an aspect ratio of k
and a taper ratio of 0.5 have been conducted. .The midchord line of the
wing was normal to the air stream. Results of these tests have been
reported in references 1 through 4. Results of tests at transonic speeds
of a wing of identical plan form and similar profile have been reported in
reference ~.
The purpose of the present report is to summarize the wing data in
terms of the static longitudinal+tability and-control characteristics
throughout the subsonic speed range of a h~othetical airplane employing
this wing. The airplane was represented b$ a semispan model comprising
the wing, a slender pointed fuselage, and a horizontal tail geometrically
similar to the wing. Force and moment characteristics of the wing, of the
wing-fuselage cortibination,
and of the complete model with two different
tail heights are presented forhiach numbers Up to 0.95 and a Reymolds
number of 2,000,000. With the flaps on the wing deflected for maximum
lift, simils,rdata are presented for a Mach number of 0.20 and Reynolds
numbers up to 10,000,000. The dynamic pressure at the horizontal tail
and the location of the wing wake are presented for the wi~-fuselage
conibinationfor the same ranges of Reynolds nuziberand Mach nuniber. The
tests of the wing-tail-fuselage combinations were conducted with various
horizontal+tabilizer settings to investigate the longitudinal oontrol
afforded by an all+novable horizontal tail. Data for an identical horizontal tail with a oonstantihord elevator (reference 6) have been used
with the wing+usel..agedata to calculate the longitudi~l+ontrol
characteristics of the model with a fixed stabilizer and an elevator.
-,----.
$%‘“-””
NFID*kNii&z
A
.<
t
.—
m
.
3
NACA RMA9101
The effective downwaah angle at the tail, the Mach number at the tail,
and the tail efficiency factor are presented herein.
COEFFICIENTS AI’lD
SYMBOLS
The following coefficients are used in this report:
%
lift coefficient lift
() ($s
CD
drag coefficient *
cm
pitching+noment coefficient about an axis normal to the plane of
symmetry ~ssing through the quarter point of tie wing mean
()
aerodynamic chord
NI
T
pitching moment
qsc t
(
)
()
Ho=
total-pressure-loss coefficient —
9.
The followin& s@ols
are used in this report:
A
a
8peed of sound, feet per second
b
twice the span of the semispanwing, feet
c
local wing chord, feet
~!
wing mean aerodynamic chord, chord through centroid of the
‘;
b/2
c2dy
f
wing semispan plan form
b/z
H
C dy
Fo
local stagnation pressure in the region of the horizontal tail,
pounds per square foot.
Ho
free+tream stagnation pressure, pounds per square foot
it
angle of tb stabilizer setting with respect to the wing+hord
plane, degrees
-Lt
tail length, di8tance from quarter point of the wing mean
aerodynamic chord to the quarter point of the horizontal–
tail mean aerodynamic chord, feet
()
.
.
N4CA RM A9101
Mach number (V/a)
Mach number at the position c-orrespondingto the centroid of
the semitail area
normal-acceleration factor of the airplane
free-stream dyn&Lo pressure
(&),
pounds per square foot
.
dynamic pressure at the position corresponding to the centroid
of the semitail area, pounds per sq~re foot
Reynolds nuniber
pvc ‘
() v
area of the semispan wing, square feet
area of the horizontal semitail, square feet
local airspeed in the.tunnel-floor bou@ary
second
.—
layer, feet per
airspeed, feet per second
—
distance from the plane of symmetry, feet
n
effective angle of attack of the horizontal tall, degrees
J-
. .
angle of attack of the win&chord plane, degrees
tunnel-wall boundary-layer thickness, inches
5
displacement thickness of the boundary layer
[f 1
inches
(1+/V)dy
1
,
elevator deflection, measurd In a plane perpendicular to the
elevator hinge axis, positive downward, degrees
trailing%dge flap deflection, measured in a plane perpendicular
to the flap hinge axis, posi,tivedownward, degrees
leading-edge flap [email protected], measured in a plane perpendicular
to the flap hinge axis, positive downward, degrees
effective average angle of downwash, positive when the air is
deflected downward, degrees
—
.
.
efficiency of the horizontal tail
visco%ity of air, slugs per foot-second
.
—
NKA
RMA9101
P
mass density of air, slugs per cubic foot
5
MODEL AND Al?PARATUS
The tests were conducted in the Ames 12-foot pressure wind tunnel,
which is a closed-throat variabltiensity wind tunnel with a lowturbulence level closely approximating that of free air.
The steel semispan model wing used for this investigation was the one
used in the tests reported in reference 1 and represented a wing of asyect
ratio 4 and taper ratio O.~.
The midchord line of the wing was perpe~
dicular to the plane gf symmetry. The wing profile was a faired double
wedge having a thiclmess+hord ratio of 0.042. The horizontal tail was
identical in plan fomand
profile to the wing and had an area equal to
one quarter of the wing area. Dimensions of the semifusehge and its
location with res~ect to the wing are given in figure 1. The semi–
fuselage was fitted tightly to the wing and tail without fill.etsat the
intersections. For a portion of the tests, tie rear part of the fuselage
was modified as shown in figures l(h) and 2(c) to study the effects of
such a modification on the yitching+aoment characteristics of the model.
The wing was equipped with a full-span, constan+chord, leadingedge plain flap and a 60.9-percen&~n,
constantihord, trailing-edge
plain flap. The area of the leading-edge flap was 17 percent of the
total area of the semispan wing and that of the trailing-edge flq was
12 percent of the total srea of the semispan wing. The unsealed gaps
between the flaps and the wing were 0.015 inch with the flaps unreflected,
The horizontal tail was mounted in the extended wing+hord plane
(figs. l(a) and 2(a) ) and alternately 13 inches (O.6g6c’) above the
extended wing-chord plane (figs. l(b) and 2(b) ). To mount the tail
above the fuselage, a bracket with a fairing body to enclose the fittings
at the point of attachment of the tail surface was added to the fuselage.
With the tail mounted in either yosition, provision was made to vary the
angle of the stabilizer by pivoting it about its >percen~hord
Ilne.
As shown in figure 2, the, semis-panmodel was mounted with the wing
perpendicular to the floor which served as a reflection plane. me gap
between the model and the tunnel “floorwas maintained %etween 0.010 inch
and 0.150 inch. No attempt was made to remove the tunnel-floor boundary
layer which, at the location of the model, had a displacement thickness
5* of 0.5 inch. The velocity characteristics of the wing+Flselage wake
at the longitudinal location of the horizontal tail were measured with a
relceconsisting of 61 total~ressure tubes and 3 static-pressure tubes.
The rake yas mounted from the tunnel floor with the total-pressure tubes
at a position corresponding to the centroid of the semitail area.
6
NACA RMA9101
.
CORRECTIONS TO DATA
c
The data have been corrected for the ef’fectsof tunnel-wall
interference, of constriction due to the tunnel walls, and of modelsupport tare forces. The method of reference 7 was used in oomputdng
the corrections to the data for tunnel+rall interference. The following
corrections were added:
&l = 0.363 CL
MD
= o.oo56~2
--
ACm=O
Corrections to the data for the constriction effeots of the tunnel
walls have been evaluated by the method of reference 8. The tignitudes
of these corrections as applied to Mach number and to dynamic pressure
(measured with the tunnel empty) are illustrated by the following table:
I
Corrected
Mach number
Uncorrected Mach
number
I
q~orrected
quncorrected
L
Wing alone Wing and
Wing alone I Wing and
fuselage I
fuselage
I
I
0.95
.92
.90
.85
.80
70
.50
.20
●
0.937
.915
; 8&7
“
799
700
.300
.200
●
●
0.917
.896
.881
.838
●7Z
.696
kgg
200
.
●
I
L 005
1.003
1.002
1.002
1.001
1.001
1.001
1.001
L 036
1.027
1.023
1.016
1.OM
1.008
1.005
1.005
The theoretical choking l.kchnuniberfor the wi~-fusel.age conibination
was 0.96.
Tare corrections due to the air foroes exerted on the turntable were
obtained from foroe meam.trementsmade with the model removed from the
tunnel.
Possible interference effects between the model and the turntable
were not evaluated. !lhemagnitude of the measured tare+rag coefficient,
based on the wing area, was independent of Mach nuniberand varied with
Reynolds nmiber as follows:
. -.
.
NACA RM A9101
7
+
Reynolds number
CD_re
2,000,000
6,000,000
10,000,000
0.0063
.0057
.0056
The rake of total-pressure tubes and static-pressure tubes used to measure
the dynamio pressure at the horizontal tail was calibrated throughout the
complete range of Mach numbers, of Reynolds numbers, and of angles of
attack of the rake.
TESTS
Lift, drag, and pitching+noment data have been obtained for the model
and its components in the following conibinations: (1) the wing alone;
(2) the wing and the fuselage; (3) the wing, the fus”elage,and the tail
mounted in the extended wing-chord plane; (4) the wing, the fuselage, and
the supporting bracket for mounting the tail above the fuselage; and (5)
the wing, the fuselage, and the tail mounted above the fuse-e.
At a Reynolds number of 2,000,000 the model was tested at Mach
numbers from 0.20 to 0.95. The range of angles of attack for these tests
was from -6° to beyond the stall, exce~t at the higher Mach numbbrs where
the range was reduced by the limitations of wind-tunnel power and of
model strength. At a Mach number of 0.20 the effect of l+ading+dge and
trailing-edge flap deflection (bn = 30° and bf = 50°) was investigated
at Reynolds numibersof 2,000,000, 6,000,000, and 10,000,OOO. ~S
combination of flap deflections was selected upon the basis of reference
2 wherein it waa shown to be the optimum for maximum lift of the wing
alone.
To determine the longitudinal control which would be provided by an
all+novable stabilizer, the model was tested with the angle of the
stabilizer varied in2° increments from-lo” to 4° for the m~elwfth
the
tail mounted in the extended wing+hord plane and from -6° to 4° for the
model with the tail mounted above the fusekge.
The velocity distribution in the wing-fuselage wake was investigated
at a position corresponding longitudinally to the midchord of the horizontal tail (3.508 wing mean-aerodynamic chord behind the quarter point of
the wing mean aero@mmic
chord) and corresponding laterally to the location
of the mean aerot@amic chord of the tail (0.428 wing mean aerodynamic
chord from the plane of symmetry). The extent of the survey was sufficient
to permit the determination of the dynsmic pressure at either position of
the horizontal tail for a range of angle of attack, of Mach number, and of
Reynolds numiber.
8
NACA RM A9101
.
An index of the figures presenting the results of this investigation
is given in the appendix.
7
RESULl13AND DISCUSSION
Force and Moment Characteristics
The lift, drag, and pitching+mment characteristics of the nmdel
and its components are presented in figures 3 through 26.
.-
Wing alone.- The effects of Reynolds number and of Mach number on
the lift, drag, and pitching+noment characteristics of the wing have been
reported in reference 1. Data from that reference for a Re~olds nuriber
of 2,000,000 at Mach numbers from 0.20 to 0.94 are reproduced herein in
figure 3. The data of this figure indicate no large or erratic effects
of compressibility up to a Mach number of 0.94. The wing lift-curve
slope was 0.062 at a Mach nuuiierof 0.20 and.increased to 0.095 at a Mach
nuniberof 0.94. The total movement of the aerodynamic center at zero
lift was only about 7 peroeht of the wing mean aerodynamic chord over the
test Mach number range.
The force and moment characteristics of the wing with various
combinations of leading-edge and trailing+dge flap deflections have been
reported in reference 2. The data of this reference indicate that a
leading+dge flap deflection of 30° and a trailing+dge flay deflection
of 50° were optimum for maximum lift. Data obtained with this combination
of flap deflections are presented herein in figure 4 for a Mach numiberof
0.20 and Reynolds numbers from 3,000,000 to 10,000,000. These data show
that deflection of the flaps increased the msximum lift of t@e wing from
0.76 to 1.4o and that the aerodynamic characteristics of the wing with
the flaps deflected were little affected by increase of Reynolds number to ‘
10,000,000.
c
The variation with angle of attack of the lift coefficient of the
wing with the gaps sealed and faired is presented in figure 5 for a Reynolds
number of 1,000,000 for &ch numbers up to 0.94. Since the wing and tail
were geometrically similar and the mean aerodynamic chord of the tail was
one-half that of the wing, these data may be.considered to represent the
lift characteristics“ofthe isolated tall and may be applied as the characte~
istics of the tail on the model at a Reynolds nuuiberof 2,000,000, based
on the wing mean aerodynamic chord, if corrections are made for the downwash and reduction in the dynamic pressure”at the tail.
The force and moment cha~cteristics of
Wing-fuselage .conibination.–
the wing-fuselage combination with the flaps neutral are shown in figures
6, 7, and 8. Comparison of these data with those of figure 3 reveals that
addition of the fuselage caused an increase in the dragj a reduction in.
.-
,
.:
...=
RACA RMA9101
9
the maximum l.if
t at Mach numbers less than 0.80, and a forward movement
of the aerodynamic center at low lift coefficients. The lift, drag, and
pitching+noment characteristics of the wing-fuselage combination with the
wing flap deflected are presented in figure 9. Comparison of these data
with those of figure 4 indicates that the addition of the fuselage caused
a decrease in the maximuu lift coefficient from 1.40 to 1.34 and an
increase of 1° in tie angle of attack for zero lift. The characteristics
of the winf-fusehge combination were little affected by a change in
Reynolds number from 6,000,000 to 10,000,000, but .anincrease from
2,000,000 to 6,cKK),000resulted in a sizable decrease in the *g.
wing, fuselage, and horizontal tail in the extended wing-chord plane.Liftj drag, and pitching+noment characteristics of the cmnplete semispan
model with the horizontal tail mounted in the extended wing-chord plane are
presented in figures 10, 11, and 12 for Mach numbers up to 0.95 and
stabilizer angle settings from 4° to —10~. At a Mach number of 0.20, the
aerodynamic center was shifted from 14 percent to 41 percent of the wing
mean aerodynamic chord due to the addition of the tail. (See fig. 12(a).)
As the Mach nurriber
was inoreased, the stabilizing effect of the horizontal
tail was diminished to the extent that at a Mach nm?iberof 0.95 the horizontal tail made little or no contribution to the stability of the model
at lift coefficients between LO.3. As will be discussed later, this
decrease in the contribution of the tail to the stability was due to an
increase in &/&L
and to a decrease in the dynamic-pressure ratio at the
tail as the &ch number was increased. With a stabilizer amglesetting
of 0° and in a range of lift coefficients of about kO.30, the complete
model was neutrally stable about the quarter point of the wing mean
aerodynamic chord at a Mach nuniberof about 0.87 and longitudinally unstable
at higher &ch numbers. At lift coefficients greater than 0.30 stability
existed at all test Mach numbers. The all-movable stabilizer provided
sufficient longitudinal control to balance the airplane model at all Mach
numbers up to 0.95 and at all angles of attack up to the stall. The value
of (%/~it)CL.O
was approximately
~.036 at a Mach number of 0.20 and
increased slightly with increasing Mach number. (See fig. X2. )
The lift, -g,
and pitching+.oment characteristics of the complete
semispan model with the wing flaps deflected are presented in figures 13,
14, and 15 for a Mach number of 0.20 and Reynolds numbers of 2,000,000,
3,CO0,000, and 10,030,000. At lift coefficients from zero to the maximum
the complete model was longitudinally stable about the quarter point of
the mean aerodynamic chord.
Wing, fuselage, and horizontal tail above the extended wing-chord
J2sQs” – To investigate the improvement in longitudinal stability and control
afforded by raising the horizontal tail above the wing wake, tests were
conducted with tie-model tail mounted 13 inches (O. 696 WinF-mean aerod~amic
chord) above the extended wing+hord plane.
10
NACA FM A9101
.
Mounting the tail above the fuselage necessitated a supporting
bracket with a streamlined body to serve as a fairing for the fittings
by which the stabilizer was attached. The force and moment cheracteristios
of the wing and fuselage with the bracket and the fairing body are presented
in figures 16, 17, 18, and 19. These data indicate no noticeable effects
of the bracket on the characteristics of the wing-fuselage combination ‘
except a slight increase in the minimum hag.
(See fig. 17.)
Lift, drag, and pitching+oment characteristics of’the complete
semlspan model with the horizontal tail mounted above the extended wingchord pkne are presented in figures 20, 21, and 22 for Mach numbers up
to 0.95 and for stabilizer settings from 4° to -@.
Comparison of the drag
data of figure 21 with those of figure 11 indicates a slight increase in
the minimum dmg which maybe attributed to the addition of the tail
bracket and the fairing body and not to the raising of the horizontal tail.
The model with the high tail was longitudinally stable at all lift
ooefflcients below the stall and at all Mach numbers, as can be seen from
figure 22. At a Mach number of 0.20, addition of the horizontal tail
shifted the aerodynamic center from 14 percent to 53 percent of the wing
mean aerodynamic chord. The contribution of the horizontal tail to the
longitudinal stability decreased with increasing Mach nuniber. As will be
discussed later, this reduction in the contribution of the tail to the
st&bility was due -primarilyto an increase in &/A
with increasing ldach
number. The ~ll+novable stabilizer retained effectiveness in longitudinal
control at all Mach nunibersand all lift coefficients.
d
—
F
There was,a marked change in the pitching+oment coefficient at zero
lift as a result of raising the tail above the fuselage. Whereas with the
tail In the extended wing+hord plane, zero pitching moment occurred at
zero lift with a stabilizer angle of 00, with the.tail raised above the
extended wing-chord plane a stabilizer setting of approxktely 2° was
required to produce zero pitching moment at zero lift. To investigate the
,
cause of this shift in the zerc+lift pitching+noment coefficient the
Reynolds nuniberwas Increased from 2,000,000 to 12,000,000 while the Mach
ntier remained 0.20. !Thisincrease had no effect on the pitching+ncment
coefficient at zero lift. Visual observation, by means of tufts, of the
flow at the afterend of the fuselage and on the tail-supporting bracket
revealed a sizable stresm angle in the region of the tail due to the rapid
convergence
of the rear end of the fuselage. This convergence was reduced
bymodifylng the afterpart of the fuselage as shown in figure l(b). The
results of tests with the modified fuselage are shown in figure 23. These
data show that, for the model with the tail mounted above the extended
wing-chord plane, modification of the fuselage caused a decrease in the
zerc+lift pitching+mne nt coefficient greater than the Increase accompanying
the raising of the tail on the otiginal fuselage.
The lift, drag, and pitching+noment characteristics of the complete
semfspan model with the high tail and the original fusehge and with the
wing flaps deflected are presented in figures 24, 25, and 26. Raising
.
.
—
.
11
NACA RMA9101
the tail a%ove the fuselage had little effect on the lift and drag of tie
model with the flays deflected. However, the model with the high tail
had more nearly linear pitching+mment characteristics than the model with
the tail in the extended wing-ohord plane.
Wing Wake and Effective Downwash
at the Horizontal Tail
The dymunic pressure at we horizontal tail, tie velocitytistiibution
in the wake of the wing-~ elage cabination, the effecti~e angles ~of’d~wash at the horizontal tail, and the tail efficiency factors are presented
in figures 27 through 36.
Looation of the wiw wake .- The location of the point of maximum totalpressure loss and the wake boundaries have been determined from measurements of the stagnation pressure behind the wiu-f uselage combination at a
position corresponding longitudinally to the midchord of the horizontal
tail (3.508 wing mean aero@namic chords behind the quarter point of the
wing mean aerodynamic chord) and laterally to the mean aerodynamic chord of
the horizontal tall semispan (O.128 wing mean aerodynamic chord from the
ylane of symmetry)s The results of these measurements are Tresented in
figures 27 and 28 where the location of the wake is presented as a function
of angle of attack for various Mach numbers and Reynolds numbers. The
location of the wake is given with respect to the wing-chord plane at 0°
angle of attack. The two alternate positions of the horizontal tail are
also identified in these figures so that the location of the tail with
respect to the wing-fuselage wake can be readily detemnined.
The tail mounted in the attended wing-chord @ane was in the wake of
the wing at all test angles of attack and at all test Mach numbers. The
high tail did not enter @e wake until the angle of attack exceeded about
70 at l&ch nuuibersbelow 0.70. As the Mach number was increased above
0.70, the high tail entered the wake at progressively lower angles of
attack. With the wing flaps deflected the high tail was above the wake at
all angles of attack. (See fig. 28. )
.
At moderate to large angles of attack and at Mach numbers above 0.85,
the wing-fuselage wake was characterized by two distinct regions of hrge
total-pressure loss. These are shown in figure 29 which presents the
variation of total-pressure loss across the wake at an angle of attack of
GO and a ~ch
n~er
of 0.85. The secondary peak of total-pressure loss
is believed to be associated with separation at the ting leading edge and
usually occurred near the amgle of attack at which the aerodynamic center
of the wing moved forward. Figure 29 also indicates ‘chatthe presence of
the fuselage influenced the msgnitude and the location of the total-pressure
losses and the location of the wake boundaries.
.
12
NACA ~
,. .. . ....-
A9101
.
Dynamic-press~e ratio and Mach number at the tail.- To determine
the ratio of the dynamic pressure at the tail to the free-stream dynamic
pressure, measurements were made of the stagnation and static pressures
in.the region of the horizontal tail, The results of.these measurements
are presented in figure 30 for various free-stream Mach numbers as a
function of angle Qf attack. The dynamic-pressure ratia at the centroid
position of.the horizontal tail inthe extend6d wing-chord plane for 0°
angle of attack varied from 0.945 at a“free+tream Mach number of 0.20 to
0.865 at a free+t+ream Mach number of 0.95. Due to the symmetry of the
model about the wing+hord plane, the dynamic-pressure ratio at the tail_
mounted in the extended wing+hord plane increased with increasing or
decreasing angle of.attack, attaining a value of approximately 0.98 at
all Mach nunbers at angles of attack of L60.
r
At a Mach number of 0.95, the dynamic-pressure ratio at the centroid
position of the high horizontal tail was unity at angles of attack less
than 2.5° and less than unity at larger angles of attack, (See fig. 30(b).)
As free+tream Mach number decreased, the minimum angle of attack for
which the dynsmic pressure remained at the free-stream value increased to
~ for Mach numbers less than O.~0.
With the wing flaps deflected, the dynamic-p?essure ratio at the high
tail position was unity, and at the position of we tail in the extended
wing-chord plane it varied from approximately 0.99 at 0° angle of attack
to approximately 0.84 ~t 10° angle of attack. me effect of increasing
the Reynolds number fra 2,000,000 to 10,000,000 was to increase the
dynamic-pressure ratio,approximately 5.5 peroent at an angle of attack of
10o with less effect as the angle of attack was reduced.
.
.
—
,_=
6
~s
The Mach numbers at the tail have been computed from the wakesurvey data and are presented as functions of angle of attack for various
free+tream Mach numbers in figure 32.
Effective angles of dcwnwas’hat the tall.- The effective angles of
downwash at the horizontal tail have been computed from the moment data
and are presented as avemge values over the stabilizer angle range in
figures 33 and 34. The e@ression used for calculation of the effective
angle of downwash is as follows:
E
=
where
(~q
)a
u+‘t -
Pf%%)=
—
(*Jai.&
is the increment in pitching+uoment coefficient due to
the addition of the tail for a constant angle of attack and (~~it)a
is the stabilizer effectiveness at a constant angle of attack. This
expression does not permit the downwash due to the wing to be sepamted
from the downwash due to other components of the model, and thus the
stream angle at the horizontal tail due”to convergence of the rear end
—
.
.._
—
13
NACA BMA9101
of the fuselage is Included in the
the data.
value of the downwash computed from
Efficiency of the horizontal tail.- The tail efficiency factor
computed from the force and moment data is presented in figures
n(@)
35 and 36. The tail efficiency factor, defined as the ratio of the lift
produced by the tail in tie presence of the fuselage to the Mft poduced
by the isolated tail operating at the same Mach nuniber,was oonqmted by
means of the following expression:
where (d~/ti)t is the lift-curve slope of the isolated horizontal tail
operating at the free-stream Mach nuder of the horizontal tail (figs. 5
and 32). No attempt was made to separate the effects of dynsmic~ressure
ratio at the tail from the tail efficiency due to the possible large
The tail efficiency factor is
variation of q+/q
alon8 tie tail s~n.
-u.
presented as a function Ef Mach number in figure 35. For either position
of the tail with the flaw neutral, the tail efficiency factor was less
than 80 percent and varied
- approximately 10 percent over the test range
of Mch nunbers and angles of attack.
.
The Effects of Compressibility
-d
The effects of compressibility on the lift, dreg, pitching moment,
in figures 37 through
and downwash of the complete model are s~rized
46.
Lift and tiaq.- !Fhevariati~n with Mach number of the angle of attack
for a constant lift coefficient tis -11
(fig. 37), increasing Mach
nuniberusually being accompanied by a decrease in the angle of attack for
a given lift coefficieq?t.
The variation with -h
nuniberof the drag coefficient for several
constant lift coefficients is shown in figure 38. At a lift coefficient
of zero, the bag coefficient of the model with the tail in the extended
wing+hord plane started to inorease at a Mach number of about 0.80. For
the model with the high tail, the drag increase started at a Mach number
of about 0.75. The Mach nuniberfor drag divergence, defined as the l.kch
number at which (~D/&)%a
= 0.10, was approximately 0.92 for the
mbdel with either tail position.
Static longitudinal stability and control.- The variation with~ch
number of the pitching-ment
coefficient for several constant lift
coefficients is showr-in figure 39. h general, the pitching+umnent
coefficient inoreased with increasing Wch nuniber. The static longitudinal
.
.
NACA RM A9101
●
instability at Mach numbers above about 0.85 of the model with the tail
in the extended win&chord plane, as mentioned previously, is evident
from the data of figure 39(a).
v
The variation with Mach nuniberof the effeotive angle of downwash
at several oonstant values of the lift coefficient is shown in figure M,
and tie variation of ~@x
wit@ Mach nuniberis shown in figure 41. For
either location of the horizontal tail, &/&
increased with .increasing
Mach mmiberbut the value of ac/~
and the rate of increase with Mach
numher was much larger for the model with the tail in the extended wimgchord plane. The static longitudinal instability at high subsonic Mach
nuniberswith the tail in the extended wingahord plane was principally
a result of this large value of a+h.
The variation with Mach number of the lift coefficient for balance
about the quarter point of the wing mean aerodynamic chord is presented
in figure 42 for various angles of stabilizer setting. The model with
the tail in the extended wing-chord plane was neutrally
stable at a Mach
nurber of 0.86 and unstable at higher Mach nuniberswhen the stabilizer
setting was Oo. With a stabilizer setting of -1° or -o the model was
longitudinally stable, but the lift coefficient for balance varied
erratically with Mach nuniberat Mach numbers above about O.~0.
With the tail mounted above the extended wing-chord plane, the
model possessed static longitudinal stability at all stabilizer settings
and all Mach nunibers. For positive values of lift coefficient, the
balanced lift coefficient for a given eJtdxLlizerangle increased as the
Mach number was increased to about 0.90 and decreased with further
increase in the Mach nuniber.
.L. -
For the model with either position of the horizontal tail, the allmovable stabilizer required only 4° to 60 of deflection to balance the
nmdel at the stall with the flaps up.
The experimental results
of this investigation have been used to
‘
predict tie static longitudinal~tability and-control characteristics of
a @pthetical airplane with a wing loading of 100 pounds per square foot
in flight at an altitude of 10,000 feet. The airplane center of gravity
has been assumed to be on an axis perpendicular to the plane of symmet~
pass,ingthrough the quarter poititof the wing mean aerodynamic chord.
The variation of airplane lift coefficient with Mach nuder for several
values of normal-acceleration factor is presented in figure 43. The
calculated effects of fligh~th
curvature on the flow at the tail were
negligible for the assuniedflight condition.
The variation with Mach nuniberof the stabilizer angle required to
balance the airplane is shown in f@ure 44 for severialvalues of nor’malacceleration factor. With the horizontal tail in the extended wi~hord
plane, the airplane would be longitudinally unstable with a normal-
—
15
acceleration faotor of unity at Mach numbers above about 0.87. Below
this Maoh nuniber,the variation of stabilizer angle tith speed was stable
and a total change of stabilizer angle of 1.P would be neoessary to
balance the airplane in level flight between Mach nunbers of O.n and
0.87.
With the tail mounted above the extended.wing-ohord plane, the
airplane would possess statio longitudinal stability at all Mach numbers
but the variation of stabilizer angle with velooity would be unstable
at &oh nuxibersabove about 0.90. A change of 2.4° in the stabilizer
anglewould be requiredto balancethe airplanein levelfli@t between
Mach nuuibers
of 0.50 and 0.95.
To compare the longitudinal mntrol afforded by the all+novable
stabilizer with that whfoh oould be accomplished with a fixed stabilfzer
and an elevator, elevato~f femotivenessdata from referenoe 6 were
applied to the hypothetical airplane. The tail model of reference 6
was equipped with a 2C@percent area, constan&ohord elevator and the
plan form and profile were identical with those of the horizontal tail
investigated herein. me elevato~ff activeness data of reference 6
are reproduced herein in figure 45 and in application of “the data it
was assumed that there was no effect of scale between Reynolds nmibers
of 2,000,000 and 1,000,000 and that the elevator efficiency factor was
100 percent.
The variation with Mach number of the elevator deflection required
to balance the airplane at the previously assumed flight conditions is
presented in figure 46.
The calculated static lon@tudinal stability and control of the
airphne with a fixed stabilizer and an elevator are similar to those
previously discussed for the airplane with the all+aovable stabilizer.
About 5&percent greater deflection would be required of the elevator
to produce the same balanoe lift coefficient as the all+ovable stabilizer.
Longitudinal Characteristics with the Flaps Defleoted
The variation with ‘Mft coefficient of the stabil.lzerangle required
to balance the model with the flaps deflected is presented in figure 47
for the model with the horizontal tail in the extended wing-ohord plane.
The corresponding &’ag ooeffioient is shown in the same figure and the
liftArag ratio as a function of lift ooefficien~ for balance is shown
in figure 48.
.
These e~erimental results have been used to preallottie powe=f
gliding speed and sinking speed at sea level of a hypothetical airplane
wiiiha wing loading of 100 pounds per square foot. The effects of the
proximity of the ground and the increased drag due to landing gear have
16
NACA RM A9101
.
been neglected. The results of these calculations are presented in
figure 49. The minimum powe~ff
sinking speed was k6 feet per second
and occurred at a forward speed of In miles per hour.
SDMMARY
r
OF RESULTS
The results of wfnd-tunnel tests at Mach nunibersup to 0.95 of a
semispan model of a ~othetical
supersonic airplane with the horizontal
tall mounted alternately in the extended wi~hord
plaqe and 0.6g6 of
the wing mean aerodynamic chord above the extended wing+hord plane have
been presented. A summary of these results follows:
.
1.
At a lift coefficientof zero, the Mach number for drag
divergence was about 0.92. There was a smooth increase of lift+.cmrve
slope with increasing Mach number up to a Maoh nuniberof 0.95.
2. The contribution of the horizontal tail to the statio longitudinal stability decreased with increasing Mach nuniber. lMis decrease
was due primarily to the increase with increasing Mach numiberin the
rate of change with angle of attack of the effective angle of downwash
at the tail. With the horizontal tail in the extended wh&chord
plane,
a further destabilizing effect was the decrease in dynamic-pressure ratio
at the tail with increasing Mach number.
3. With the horizontal tail in the extended wing-chord plane, the
model was longitudinally unstable at Mach nunibers above 0.87 at lift
coefficients less *n
0.3. With the horizontal tail 0.696 of the wing
mean aerodynamic choziiabove the extended whg-herd
plane, the model
was longitudinally stable at all lift coefficients for all.Mach numbers
for which data were obtained.
4. Eitier anall~vable
stabi~zer or a f~ed stabilize rwi~ a
constan-herd elevator provided sufficient longitudinal control to
balance the model throughout the test range of &ch numbers.
Ames Aeronautical Labomtory,
National Advisory Committee for Aeronautics,
Moffett Field, Calif.
.
,
‘
17
NACA RM A9101
.
APPmimX
The follmd.ngtableshave been includedto providea convenientink
K
to the figures
presentingthe resultsof this investigation:
FORCEAED MQ4RTT ~TICS
wing
Results
presented
a,@,
&Cmvz
Flap
defl.kction
C~
00
a,~,&hVSCL
Alone
Reynolds
Mach
number
0.20
to
2X106
3
3xlo*to lo%lo=
4
0.94
0.20
on=30°,*-500
Figurenumber
number
W- Alone With All.Gaps Sealed
Results
presented
Flap
deflection
---
CL Ys a
Mach
number
0.20
I
to
0.94
Reynolds
number
Figurenumber
Moe
5
W~Fuaelage Cabination
.
.
*
FIEP
I deflmtinn
Remits
presented
I
00
CLvsa
0.20
1
CLVSQ
CL W CHL
on=300,&=500
a,q)&cmvscL
Results
presented
Mach
number
map
deflectlml
to 0.95
I
Reynolds
number
I
Figurenuniber
6
2X1CP
7
I
i
8
0.20
2xlo= to lmlC@
9
Stabillz=
angle
number
40 to -MO
0.20
Figurenumb=
cLvsa
CL PE ~
CLVZ~
CL~a
CL
VE
5n=30°,%-500
2X1(Y to lmlo=
~
cLvsc~
1
i
I
I
is(a)
to 13(C)
14(8)
to M(c)
~(a)tou(c)
NACA RMA9101
WiwFuEelage
Combination with Bracket for Mounting Tail Above Fuselage
r
FlaP
deflection
Results
presented
Reynolds
Mach
number
0.20to 0.95
00
Figure nmber
number
I
>
—
2X1O=
I
16
17
18
I Woe
0.20
Ei&300,5f=500
Wing, Fuaelag%and Horizontal Tail Above Extended Wi@hord
Re6ult’a
preaepted
CLva
F1.q
deflection
a
00
Mach
number
Stabilizer
angle
40 to-60
Reynolds
number
2)(106
2.20to 0.9!
19
KDaoe
to
Plane
Figure number
h
20(4
4
to 20(h)
CL V8 CD
21(8) to 21(h)
CL VS Cm
22(a) to22(h)
%Lva
00
cm
cL Vs a
0.20,
0.92,
40,00,&-W’
&=300,6f=500
23
acloe
0.90
0.93
ZRIIYto 10X1OC
0.20
24
cL
va CD
2,5
CL
V6 cm
26
Slmwathe effectof modifyingthe rearof the fuselage.
FLOW CONDITIONS IN THE REGION OF TEE ECRI,ZONTALTAIL
Chmacteriatica of Win&?uaekge
Results
presented
Location of
va a
wake
FlaP
deflection
o“
Mach
number
Wake
Reynolds
number
0.20to 0.95
2KL06
Figure
number
27(E4t027(g)
I
Z$l=300,5f==oo
Pressure loaa In
wake vs distance
from win~hord
plane
00
0.20
0.85
2X1O=
to loxlo~
2X106
28
29
I
.
19
NACA RM A9101
~LT
d
R-=tio, Meeh Shmhr,
~=e
m
deflection
aeto
0.20
the
to 0.9s
mil
Hgum
mmber
30(a)
& 30(b)
31
llxuoe
2%UP
32(w)& 32(b)
0.20 to O.*
2xLOe
33(a)& 33(b)
0.20
2x10’3to lmloe
34(E)& Sk(b)
0.20 to o.g5
2XKP
35(.) & 35(b)
00
0.20
00
8U=@,8#30
00
at
we
toO.*
0.20
8n-3@,8#
DOlumaeh
ReymJI&
number
HELoh
00
of
2xfe0tiv0
Angb
0.20
%=3dhf-60°
36(a) & 36(b)
2uo=tolo%loe
sLM4#rtY CORVE!3
Zffeots of Cmuressibility cm the Chareoterietfcsof the Mcdel
de~lectfon,d;
Reynoldn Iu&ber, W]
The
[KL8p
Lift emidrq
{
Reeults
Idft
Coefficient
me8ented
ava~
O to
~TaM
Stubillmr
engle
Maoh
number
0.6
J
me
numbcm
0.20to o.g5
37(a)& 37(b)
i
38(E)& 38(b)
r
stnbtli~ EuMiIxmkYl ehuaateristical:
Lai8itudinel
.*
Resultm
Stabilizer
I&t
caefficklt
prenmted
.’
o
Meoh
number
0
0.6
3g(a)
b(a)
~
41
T
00 to -r
---
I
---
[
0.50to
-----I -----
---
I
it
for
~.~
qp3
Klap
deflection
&
0.20
---
Mlgitudinal ChareotertitloaWIReeults
preeentea
Mach
number
requir-tm ofby@aletiodeirplme With
au-
d
0.20b O.gk
ks
0.20
h’6
to
o.g5
nmbac
a
b7
wing
UaMng
“1 :
of103pmvMs
perOquere footin
lo,wJ feet.
npeedfor bypothetiod airp@ae
vith
flight
at neaI.mel.Poueroff.
%Mnklng
4.4
lctdo=
Kpeed m
●med
~
o.g5
R=-
0.2U
‘-’111!
at
43
,q+o
7
zdft
-t
to
k?(a) & &(b)
O.*
the r-ley Defleotea
L/l)Y# CL for
%nklng
miung
~g(b)
k b(b)
&
a
wing
I.oeMng of
loopomde~B-foOt”in
20
NACA RM A9101
..
REFERENCES
1.
Johnson, Ben H., Jr.: Investigation of a Thin Wing of Aspect Ratio
4 in the Ames 12~oot Pressure Wind Tunnel. I - Characteristics
of a Plain Wing. I?ACARMA8D07, 1948.
2.
Johnson, BenH., Jr., and Bandettini, Angelo: Investigation of a
Thin Wing of Aspect Ratio 4 in the Amss lE%Foot Pressure Wind
Tunnel. II -The Effect of Constant-Chord Leadi~
and Traili~
Edge Flaps on the Low~peed Characteristics of the Wing. NACA
RMA8F15, 1948.
3*
Johnson, BenH., Jr., and Demele, Fred A.: Investigation of a Thin
Wing of Aspect Ratio 4 in the Ames l~oot
Pressure Wind Tunnel.
111- The Effectiveness of a Constan~hord Aileron. NACA RM
A8117, @l.8.
4.
Johnson, Ben H., Jr., and Reed, Verlin D.: hvestigatfon of a Thin
Wing of Aspect Ratio 4 in%he -s
l~oot
EtressureWind !lhmnel.
rv– The Effect of a Constant<hord Ieadin@!ldge Flap at High
Subsonic Speeds. NAcARMA8K19,
1949.
5*
Rathert, George A., Hanson, Cal M., and Rolls, L. Stewart: Investigation of a Thin Straight Wing of Aspect Ratio 4bythe NhCAW~
F1OW Method. -Lift and Pltchin@40ment Characteristics of the Wing
1949.
Alone. NACARMA8L20,
6.
Bandettini, Angelo, and Reed, Verlin D.: The Aerodynamic Characteristics Throughout the Subsonic Speed Range of a Thin, Sharp+l!klged
Horizontal Tall of Aspect Ratfo h Equipped with a Constant-Chord
Elevator. NACARMA9E05, 1949.
Sivells, James C., and Deters, Owen J.: Jet~oundary and PMorm
Corrections for Partial-Span Models with Reflection Plane, End
Plate, or No End Plate in a Closed Circular Wind Tunnel. NACA
Rep. 843, 1946.
Herriot, JohnG.:
Blockage Corrections for ThreeAlime~ional-low
Closed-Throat Wind Tunnels, with Consideration of the Effect of
Compressibility. ItACARMA7B28, 1947.
.
k
—
Af/ dimensions given h
jnches unless otherwim
specifiti.
.—.
—
. +.
ffote:hwdi~edge radii are
i2
and trai7hg0.005
inch
;*
—-~-—
Sym
~-Wing and tail section
‘!
&nvdiJ
(Fan=#lewi#,
Fus.abLw
36.00
(u)
Figure
l.- Semi-
horizontal
modsi
fal/mwmfedin
of on iw~bns
extended
wing-chord
with a wtng and all-
plane,
movabk
horizontal
tail of a~cf
ratb
4.
I-i)
P
NACA RM A9101
22
All dimensions
given in inches
unless otherwise
specified
.,
horizontal
toil above the fuselage
~ of horizontal
Original
fusehge,
--------
.//,
R
toil
Modified fuseluge 7
I
\
------ -7-:=’--+:..
,
-1
●. .
._.
Modified fuse/oge
coordinates perce~~ length of
“{ originu/ fuselage )
7
L
Toil foiring body
coordinates
[percent length)
r,
x,
7.%
/4.8/
22.22
29.63
37.04
50.00
62,96
70.37
?7.78
95./9
92.59
0000
4.$6
/4. 58
/8, 48
2Q 78
21.85
22.22
2/.85
20.78
/8.48
/4. 58
4.96
(b)Hori;ontol
/. - Concluded.
r=
0
toil mounted
moolfieo’ fuseloge.
Figure
x=
425
6 5
i97:
7:9;
3.960
8220
8$?09 3.890
3.822
9590
/02.80 2897
2.040
/0825
10S?60 /.746
///.00 /.438
112.30 /./03
//3.70 0.758
above the original
un~ the
1“
,
(a)
Figure
2.- Semlsp+m
model
Horizontal
tall nmnted
of the airplane
nmnted
In the extended
3n the AJIBS M-foot
wing-chord
pressure
plane.
witi
tmmd.
.—.. .-.
.
.—.
—.—
#
... ,.,
(b)
F@ur(j
2.- Conti.wd.
Eorizontil
,,,
tail mmnted
,,
Fdd1::,
above
the I%ebg..
~k-. .- .-d
.
.
I“
i
i;’
‘,
(C)
mgure
2.-
continued.
Modified
fueelage
with the horizontal
tail munted
above
the ~elage.
.
.
.
E!
(d)
F@ure
2.- Conoluded.
Ralw of premsura tubes muntedkmhirtt
hew~uselage
conbinatl.on at the location of tho horizontal tail.
.
.
.
-.
,
●
r
b
ii!
Ifvldl
I
I
I
I
I
v b~
d
-,4
-.6 E
1-
.fjj_
04
~
Figure
3.-
D8
./2
for”iW,
.16
coeftkien~
0.20
-4
0
4
8
Angle of offock,a,
CD
71e lift, drag,ond
-8
pdching-nwment
churocteriktics
/2
for
M,
deg
of
the ploin
0.20
./
o
+
PitcMg-moment
wing. /?, 2,000,000.
forluo20
coa%bieti,
-
‘
Cm
.
c+
.
s
*
$
\
9)
o
Q
Q
..
-1
0
.04
a08
Dtvg
./2
.16
coet%kw~
.20
.24
CD
.28
.32
-8
.36
.40
of
otfmk,
-4
Angh
Figure 4.- Tbe lift, drag, and pitching-moment
M, 0.20.
choructeristics
0
.44
4
8
12
16
a, deg
of the plain
20
J
wing with fh
o
+
-2
RMJkwm@
flops deflected
-IW
&, 30°;
“,3
L
+,
50-;
g
H
o
P
*
,
Q.
I
,,
,, .
1
NACARM
A9101
33
/.2
M
/.0
O
0.29
❑
0.50
0.70
0.80
0.85
0.90
0.9/
0.94
0
A
.8
V
v
~
.6
.
v
}
c
-.4
n
h
d
-.6
‘.8
-/. o
.
-1.2
-/6
-/2
-8
-4
0
Angle of
Figure 5.- The lift characteristics
attack,
4
8
a,
/2
for
A4=0.29
deg
of the plain wing, gaps sealed. R, ~000,000.
34
NACA RM A9101
.
●
.
‘--8
Figure 6,–
-4
0
Angle of uttuc~
The lift
chorucferistlcs
4
8
/2
/6 for M=O.20
a, deg
of
the wing
and
fuselage.
/?l 2,000,000.
.
NACA RM A9101
35
...... . . . .
—
Figure Z— The dreg characterisflcs
of the wing und fuselage.
R, 2,00~000.
NACARM
36
A9101
.
/.2
M
o
/,
.
0 0.20
13 0.50
00.70
~ 0.80
v 0.86
B 0,90
Q 0.93
F 0.95
.8
.6
.4
us’
“
2
.
v
-.4
-.6
,/
-,/
o
Pitching-moment
Figure 8+
fuselage,
The pitching
R, 2,000,000.
for
coefficient’,
-moment
M=O.20
T
Cm
chorocterlstlcs
of the wing and
.
,
E!
I!!
mzqivehwm,~
-8
-4
An@
0
of
4
ott@k,
8
@
16
20
a, d~
71
Ru
Figww
9,-
+, 50”;
The Iift,drag,and
M,o.20.
pitching-moment
choractetisiics
of the wing and fusebge
wi?hthe flaps
.2.
WemYk@
deflected
f%!
an, .30”;
....
-Q/#
38
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A9101
.
.
v -4
*
1/
.
!
WwFFFFm
1
,
!
m
m
I
I
I
.&?
-8
-4
0
&g/e
4
1
I I I {
/2
8
16
t I
20
I
for
I
I
& = 4°
Of 0ttc7Ck, a,deg
(U)M, o. Za
Figure
/0,–rhe
lift
chart?ctert’sties
mounted in f~e extended
wing-chord
of the airplune
plane.
model
R, 2,000,000.
with the horizontal
foil
1
39
NACA RM A9101
/.0
.8
.6
.4
H
H
00
A ‘2
‘-ol%ir
D -G
4
-8
v -/0
LHii5EE
u-LK
-.4 L
v
JT
1/
YY
-.6
[
I
1
r
I
1
1
1
1
P
1
t
-.8
-8
0
Angle of attack,
-4
48
u,deg
/2
(b) M, O.50.
Figure
/O.-
Continued.
/6
for ~= 4“
TYH
40
NACARM
A9101
.
.
*
N*
.?
Q
$
Q
u
$
4
-8
-4
4
Angle of attack,
0
8
a,09g
/2
/6
for 4 “4”
(c) M, 0.70.
Figure
/O.-
Continued.
.
.-
NACARM
A9101
41
.
.
/.2
/.0
.8
.6
.
.
-.2
1/
-.4
1A
141
I
I
t
I
I
I
I
I
.=
-8
-4
Angle of
0
affack,
4
8
/2
a,o’eg
ld)
Figure
10.–
GOnI%7U@d.
16
M,
0.80. ‘
20
for ~ “ 4°
I
i
I
.
42
NACA RM A9101
.
.
/.2
+-
/.o
8
.6
A ‘2
l-i
v -4
II
v
4I
-/0
,4
*
/1 /1
v
F‘
/
iv
P
?
.“
-8
-4
0
Angle of attack,
I
8
I
-=zi=-
/2
/6
for ~= 4“
u,deg
(e) M, 0.85
Figure
/0.–
Continueo’I
.
.
NACA RM A9101
/.2
i
3
/
/.0
“n
n
L1
u
04
.6
I/
A
1
I /
H
C12
/-
/
m
/1
00
.4
H
V-4
.2
0
-.2
I
I
.,
Y
1/
-f.o
-8
–4
‘4
0
Angle of attock, u, tfeg
8
/2
[f) M, 0.90.
Figure
1o.-- Continued.
for ~ =4”
,
Figure
/0.-
Continued.
.
45
NACA RMA9101
.
*.
b
04
t--l
.4 u
.2
A
‘Z
v
-4
D
-6
H
I I
—
..-
v -/0
,
I
I
‘.8~
-4
‘8
Angle
0
of dtuck,”a,
I
I
4
1
L_
8
deg
(h) M, 0.95.
Figure
10.–
Concluded.
for 4=4”
46
NACA RM A9101
.
.
10
.8
.6
.4
&
~’ .2
.$
80
g
4
-.2
-.4
.
—
-.6
-.8
0
.04
.08
.12
.16
Dog
.20
.24
coefficient,
.28
.32
for ~ ‘ 4°
Co
(U)M, O.20.
Figure //.–
The drug charucterisilcs
of the airplane
model
mounted in the extended whg-chord plane. R, 2, QUO,000.
c
-——.
with
the
horizontal
toil
NACARM
A9101
47
‘
.
.
/.,0
.
8
6
,4
,2
‘o
~
4
-Vttttli”
-.2
-.4
-.6
-8.-
0
.04
Drag
.08
./2
coeffkien~
C’
./6
.20
.
[b) M, 0.50.
Figure
Il.—
Continued.
.24
.28
for
4=4
4a
NACA RM A9101
.
/.0
.8
.6
.4
6’
S*
.2
.?
f.
$
4
“.2
-.4
-.6
.—
-.8
0
.04
.08
./2
Drug coefficient,
./6
.20
.24
.28
.32 for
~= 4°
6“
(C)M, O.70.
Figure
Il.–
Gontinued.
.
49
NACA RM A9101
/.2
10
.8
.6.
.4
e
~c
.: .2
.%
k
Q
Q
‘o
~
4
-.2
-.4
-.6
-.8
.04
0
.08
Drag
./2
.20
.16
coefficient
GD
(@M,
Figure
.
//.–
Continued
.
.
.24
0.80.
.28
for lj “ 4°
50
NACA BM A9101
—----
,
.
AZ
/.0
.8
.6
.4
‘~- .2
.%
fo
●
Q
v
.$ -,2
4
-.4
-.6
-.80
.04
.08
./2
./6
.20
.24
(e) M, 0.85.
Figure
//.—
Coty’inueo!
_ ........ . . . . .
.28
for
I)= 4°
51
NACA RM A9101
/.2
/.0
.8
.6
.4 !
I
Ua
V4
.-
v
\ .1
\
*
\
\
0
.04
.08
Dreg
coefficient
.20
Go
figure
II.–
Continued,
for
7
II
I)*4”
... ,
(f) M,
.
.
J6
./2
,“
0.90.
NACARM A9101
●
.
a
“
Figure
//.–
Continued.
53
NACA RM A9101
.
.
*
.
.
.04
o
.08
Dug
//.–
Goncluded.
./6
coefficient,
(h)M,
.
Figure
./2
CD
0.95.
for ifs4”
.6
.5
.4
.3
.2
./
O
:/
(WM,
12.- The pitching-
horizonkd
toil
moment
mounted
in the Wended
.
:5
:6
;7
Cm
ii!
of
wing-chord
.
-.4
0.20.
characteristics
,
:3
metficieti,
FVching+noment
Figure
:2
the
airpkme
pkme.
model
with
the
R, 2,00~000.
.
.
✌
✎
●
☛
/0
.8
,
I
.6
f’1
A
J/
r
.
~,+’
.
H++
.4
H-ii—t—H—l
.2
i
,,
!
I
i
.1
.6
.5
.4
.3
2
./
Pitching-moment
(b)M,
Figure
12.-
I
I
I
I
I
I
-.8 I
I
Continued.
o
./
-2
c~ftkien~
:3
4
?5
I
I
I
I
-6
I
i
?7
Cm
a 50.
U
U
ul
m
m
.8
.6
@
.4
g
\
.2
1
0
Ill
5
-.2
-.4
-.6
-.8
.6
.5
.4
.3
.2
J
O
Rkhing+noment
:1
coeftM@
?2
73
-.4
75
-6
~
(c) Aft 0.70.
figure
/2.–
Continued.
1
.
1
.
,
#
a
/.2
g
o
P
,8
.6
d
-.2
-.4
d
II
I
I
r
1/
r
I
-.6
w
Fiiiiiiiiiil
1
1
/
I
I
6 Wtrwt
1
I
1
I
1
1
I
I
I
I
I
1
1
1
1
I
I
I
I
I
A
1
1
tum
UJ r
1
-.8’
7/
.54.32]0
fYfdi~+xwhm
t
:2
cdfuk?n~
(d) Ai, 0.80.
figure
12Y
Continaf.
-3
G
-#
-5
76
:7
.
.-
i.2
/.0
.8
.6
e
.4
+
$
.5
Q
g
-2
o
4
-.2
-4
-.6
-B
.63”
.4.3.2./0
:1
IWcbihg-monkmf
{e) IV,
figure
.
/2.-
co&ihi”m~
72
-3
-.4
:5
-.6
5!
G
0.85.
Continued.
1
✌
●
‘
[2
/,0
.8
.6
.4
t+
,2
.
$
“$
0
%
$
-.2
-.4
-,6
-.8
-lo
...6
,5
,4
.3
.2
J
O
+
c06mim~
R7dlihg-mment
(f)
i7gure
i2,–
Continued.
M,
0.9Q
72
~
.3
+
.5
d
33
OH
I
,
IQ
I
I
1
f
I
II
I
I
I
A
I
t
,
“tb!tkki
76
/
/?
v
/
/
x
v
k
r
P
f
/
/
3
4
$
1
Jlllliil
/
-J3-
.6
.5
4
.3
.2
,/
RwVhg—mamn
(9) ~,
Rgufl
/2.–
I
I
I
1
I
I
L/
Continued.
.
o
t
-]
cv#%bkw~
0.92.
:2
G
-3
4
-5
*
.
figore
12.-
Concluded.
62
NACA RM A9101
. .—
.
. .-
.8
.6
.4
w
.2i
R
w
.
o
.s
-.2
/
/
/
-.4
=4s=’
c‘
‘.6
-8
-4
0
48
/2
Angle d
(a)R’,
Figure 13–
The lift
characteristics
mounted in the extended
dn, 3(7; ~f ,50°
; M, 0.20.
uttuck,
/6
20
for ~= 4*
a, deg
2,000,000.
of the airplane
model
wing-chord plane and
with the
horizom’ul
with the flaps deflected.
tall
..””
“.,
.:.,
NAC!ARM A9101
Angle of attack,
a,a’t?g
(’lot?, q
●
figure
.
.
13?
Continued.
000,000.
63
64
NACA RM A9101
.
II
H
u
/.0
.
%-
04
❑
12
00
II
I
.6U
Q -8
I
I _,/’f p
J-( ~/
4%4
.4
.
.2
.
0
. ‘.2
-.4
‘L
‘~L
1/
1/
y,
I@
f
T‘
/
c,
?
<(
1
-.6
-8
-4
0
4
8
Angle of ottoc~
a,deg
(C)R,
Figure
13.–
1
I
/2
I
I
I
16
10,000,000.
Concluded.
@ya?%wEIi?
-.,.-. . ....
I
I
for ~=4”
I
I
I
I
I
NACA RM A9101
65
.
●
L6
L4
12
/.0
.8
.
G’
~- .6
.=
Q
s~ .4
Q
s
.2
0
-.2
-.4
-.6
0
.04
.08
./2
.20
./6
Drag coefficien~
.24
Go
.28
.32
for 1;= 4*
(0)t?, 2,000,000.
Figure 14.– The drag characteristics
.
mounted
8f,
.
in the extended
50°; M, 0.20.
of the airplane
wing-chord plane
model with the horizom’al
with the flaps
deflected. &
30°;
tail
66
HACA RM A9101
1.(2
.8
.
.
0
.04
.08
./2
./6
.20
.24
.28
Drag coeffickm~ Go
(b) R, 6,000,000.
Figure
14-
Continued.
.32
for
t) = 4“
NACARM
67
A9101
a
.
.
0
.04
.08
.12
.~6 20
Drug coefflcien$
.24
.28
Co
(c) R, 10,000,000.
Figure
.
.
/4. - Concluded.
for~=4°
68
NACA RM A9101
-4
,
.
.
-i
?
.
.6
.5
.4
.3
.2’
./
Pi%%ing-moment
O
+
coefficiii~
:2
-3
-.4
-5
~
(0) R, 2,000,000
Figure /5. - 7he pitching-moment
chffrocteristics of the oirplune model with
the horizontal tai/ mounted in the Wended wing-chord plane and with the
f/aps deflected ~, 30~ +, 50~ M, 0.20.
-.—
.......--..
W“d
69
NACA RM A9101
.
.6
.5
4
.3
?2
.2
J
O
J
Pitching-moment co~cb~
(&
(b) R, ~000,000.
Figure
15F
Gontinued.
?3
-.4
?5
;6
NACA RM A9101
70
.
.-
.-
--.
.Q* . b
o
4
“S
❑
2
o“
-2
~
%.
4
0
.4
A
-6
4
-8
v
-/0
!
D
.2
,n . ,
/’%for.zontul
r
}1’1-l
0
.—
-4
v
El
b
.
toil off
-.2
-.61 ‘ ‘“
.6
Figure
.5
/5.–
I
I
I
I
.4
Concluded.
I
.3
I
1
I
,
I
1
1
!
1
1
72
+
.2
.1
0
Waling-moment coatficien~ C%
(c) R, /O,OO(?,000.
1
..
I
:3
-.4
75
71 “
NACA RM A9101
.
/.0
M
.8 —
n
o
o,5~
0.70
A 0.80
D 0.90
F —
a 0.92
v 0.95
.4
6
-~
c
.-a!
u
.2
c
g
.
.
.
o
g
4
.
.
T
/‘
-.2
>
,
L.
/
/
77
14
) V
/
‘.6
TP
‘<
‘F
.
-B
‘8
-4
0
4
Angle
Figure
for
‘
16.—
~he lift
choracferisiics
moum’lng the horizontal
fail
/2
8
of oftuc~ a,
of the
ubove
the
wing
/6
20
for
IW= 0.20
o’eg
and
fuselage.
fuselage
with
R, 4000,000.
?%e bracket
NACA R’MA9101
*
.
.
.
.“
.
b
Figure
the
.04
.08
,/2
./6
,20
,24
Bag coeflicien~ Co
/7. - The drffg L2$ff
rueteristics
bracket
R, 2,000,000.
for mounting
,28
for
M= 0.20
of the wing ond fuseloge
the horizontal
with
toil ubove the fuseluge.
.
—
“..
NACARM
73
A9101
/.
o
[
Illllllld
I
I
I
I
I
I
m
I
I
I
Islll
m
.8
r
/
\
-.
I&l
.w
.6
<
IY
I
I\
1 rJ
I
.4
\
\
I /1
141 d
/1
w
II
D
/
I
IY-HHI(I
I
\
I \ I
w
s
I
II
121
A
i
I
Ihf
00.20
I •10.50
\ 00.70
A 0.80
\ V 0.85
~
\I
.,
I
I
.
.
-.2
‘. 6
‘.8
-./
o
J
for
Pitching-moment coefflclent,
Figure /8.-The pitching-moment
fuselage
.
.
wlfh fh8
the fuselage,
bracket
#?t 2,000,000.
M=O.20
CM
characteristics
of the wing
for mounting the horlzontol
uno’
fuil
above
.8
.6
-.4
-.6
0
.04
.08
Dug
.E
.16
CO#w@
.~
CD
.24
.28
-8
..3?
-4
.36
0
Am
of
.
.40
4
a~ck,
8
/2
a, dq
E
20
./
o
-.1
Fh’chlng —nwmew
Figure j9.mounting
.
.
Zhe Iift,drag,
the horizontal
and pitching
toil
above
‘momwt
charoc~e~~tics
the fuselage,
.
flaps
of
deffedeo!
the m“W O@
+ , 30~
{,
fUSehI@
5fl
wifh
the
M, 0.20.
.
.
?2
co&%4m~
bracket
(&
for
E!
g
t-i
o
1-
NACA RM A9101
r
.
/.0
.8
deg
04
❑
.6
‘t
2
00
A ‘2
V4
&
.4
D
H
-6
.
.i
‘8
0
-4
Angle
of otfoc~
4
GZ,deg
(’U)M,
Figure 20.fail
~he lift
mounted
characteristics
above
the fuselage.
.
8
8
/6
/2
for tj*4”
0.20.
of the ui~lune
R, 2,000,000.
modei
with the horizontal
. .
76
a!MaEEe)
.. ..-.. .
NACA RM A910L
.
.
.
(b) M, 0.50,
Figure
20.–
Gontinued
.
.
NACARM
A9101
77
/.0
.
‘t
deg
.8 —
04
00
.6 —
A
(j
‘z
v-4
.4 —
P
‘6
/
(J
//
/
4
/
t
.2
)$
/!/
*
o
-.2
z.
\-
1
$-
F~
●
(
J(
~
<>
-.4
c)
r
L
) ~
,
c
<(
-.6
L>
Kr
{
D‘
~–
-.8
+
-4
0
4
Angle
/2
/6
8
of otfuc~ U, deg
(c) M,
Figure
20.—
0.70.
Continued.
9
for it=4*
78
NACARM A9101
.
.
/.0 -,
~,
,8 —
deg
U2
.6 “
00
A ‘2
.4 —
G
~.?
u
s
w
~
V4
.2
0
$
4
-.2
.
.
.
‘.8
‘8
+
o
4
/2
8
/6
Angle of ottac~ u, deg
(d)M,
Figure
20.—
Gontinued.
0.80.
for #j =4°
79
NACA RM A9101
‘
.
/.0
.
‘f
,
?
deg
8—
04
❑
.6 —
z
3
00
/
A ‘2
-,
c.5—-
“TtttT7m
-.8
FF#
-/.0 ~
‘8
I
–4
I I I I
o
4
Angle
of uttac~
8
/2
a, deg
(e] M, 0.85.
Figu re
.
.
20.—
Gontinued
/6
for
~=4*
80
NACARM A9101
.
/.0
.
‘t
deg
.8 —
04
R2
00
.6—
‘A ‘2
V4
.4—
D -6
..—
‘8
0
-4
Angle
of attack,
4
a,
8
deg
{f) M,
Figure
20.–
Continued.
0.90.
/2
for 4=4”
.
NACARM
81
A9101
.,
0
-4
‘8
Angle
of ufftick,
4
a,
&g
(g) M, 0.92.
,
Figure
-.
20.—
Continued.
8
for it=4”
&?
...’-...-
-
NACA RM A9101
.
—
.
.
*
Figure
20. – Concluded.
.
NACA RM A9101
83
.
.
Drug coefficien~
~
(a)M, 0.20.
Figure 2L—
The drag
of the uirpkme model wii’h the
R, 2,000,000.
chorucieristics
mounted ubove the fuseluge.
horizontal
fail
NACARMA9101
84
.
.
,
.
-.UO
.04
.08
./2
.16
N.11”1111
.20
Dreg coefficient,
(b) M, 0.50.
Figure
21.-
Continued.
~i
.24
Go
.28
A
for & = 4°
I
NACA RM A9101
(c)M,
Figure
2/.–
Continued.
0.70.
HACA FM A9101
86
/.0
I
.8
-.
.6
.4
~
-1
-.2
-.4
I\.
-.6
.-.
0
.04
.08
./2
./6
I
.20
I
I
.24
Drag coefficien~
(d)Af, 0.80.
Figure
2/.–
I
.28
I
I
fw&4°
Go
,
Continued.
.
.
NACA RM A9101
,
,
1.0
.8
/
/
.6
j
/
deg
/
/
Y
.4
/
,
?
#
/
k
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●
.
$
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4
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00
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o
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.
-.4 —
Y
—
-.6
‘.8
-/.0
o
.04
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./2
./6
Drag coefficien~
.20
Go
(c)M,
Figure 2L-
.
Gontinued.
0.85.
.24
.28 for @4°
88
NACARM
A9101
.
#
.
.“
-.
—
.
.
NACA
,
.
.
.
.
.
Dreg coefficient,
CD
(9) M, 0.92.
Figure 21.- Continued.
●
.. ..
.
NACA RM A9101
9
.
.
P
/
I
A
()
\
\
I
I
\
.
-.4
.
o
.04
Bag
.08
coefficient,
./2
G~
W
figure
2/.—
Cone/uded.
./6
M,
0.95.
for ~S4e
.
.
Ill
Pitching-moment
Figure
the
22.—
The
fuseloge,
pitching-roomed
R, 2,000,000.
characteristics
of
coefficient
{0)
M,
the
air.okw
Gm
0.20.
modei
with
the
imrizonfui
taii
mounted
.
—
1
1
+
.8
04
H
1
1
I
I
I
I
,
I
I
I
t
I
,
,
I
,
1
,
deg
H
.6 H
1
I
❑
I
I
z
00
A ‘2
,4
,,
I
1/
I
)
~i
1
I
1
Ciii
:.
i
! ,,
!.
,.
;
L
I
-.4
-.6
,
–.0
.6
.5
#
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1
t7tii@—momen
0
22?
:2
-3
+
s
s
-5
t cQIS%ficim~ l%
(b) M, 0.50.
Figure
7/
‘!2
&H“
0
I-J
Continued.
*
.,
,
.
,
,-*
!$!
UY
u
U3
Lo
.F
.8
.6
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e
~s
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~
*
lo
g
4
-.2
-.4
-.6
-.8
.5.4
.3.P
71
’JO
Pitcting-nmment
coefflcjen~
(d)hf,
figure
.
.
22.–
%?
:3
-.4
+
~
0.80.
Continued
,
,,,
.
.1
Ill!!l
,7
.6
.5
.4
.3
.2
J
FrMiw?wmw
fe)M,
Figun3
22.—
Continued.
t
7/
O
meftkkq
0.85.
72
G)
-3
-.4
75
76
II!l”l!!l!l!bll?
j
.8
H-i
I
I
o’eg
m
o
4
❑
2
I
I
I
I
I
l#T
I
<1
I
!/
8
r
II/
‘<
.4
-.6
-.8
-10
.7
.6
.5
.4
.3
.2
J
Pltchihg-nmment
(f)
Figure
.,
22,-
Conflnued.
M, 0.90
O
-./
coefficient
-.2
Cm
-3
-.4
-.5
.8
.6
.4
-.6
-.8
-10
.7
.6
,5
#
.3
.2
.1
Rhhing+mment
(g)/U,
Figure
22T
Continued.
0
cafXWn~
0.92.
‘~
‘.
&
?3
-.4
“5
.
,71
:,,
,.
.,
S6
/1
J
1/1
v l.?!
1
‘“2
I&.
.C
1,
.
I
Pitching-moment
coefficient,
I
I
I
Cm
!2!
(w IW,O.95.
Figure
I
22.-Concluded.
1
99
NACA RM A9101
.
.
M
O
❑
.8
0
A
0.20,
0.20,
0.92,
Originul
Afodifie’d
Original
fuselage
fuselage
fuselage
0.93,
Modified
fuselage
.6
f
.4
(3’
~w
~
.%
Q
$
m
Q
.
Q
~
4
>
.2
4
o
<
}$
-.2
(
/
.
P
-.4
/
{
~$
J
$
{
/
v
/
/
‘.6
+-+V-
4Q
‘.8
.3
o
./
.2
Pifchingwnoment
4
.3
-f
:2
coefficien~
Cm
.2
./
:3
o
Pitching-moment
Figure 23.- The effect on the pitching-moment
-1 ‘
coefficient
characteristics
fuselage of the airplane model with the horizontal
/?,2,000,
000; $,00.
-.4
72
:3
Gm
of modifying the
toil obove the fuselage.
100
XAC!ARM A9101
.
7,4
4
I
1
1
1
4
1
/.2
Q-
“-
rr
deg
o
d{
0
h
1
i I?
1q
-.2
*
/
f
+\”
n
2,000,000
4
D
6,000,000
4
0
/0,000,000
4
A
2,000,000
0
v
6,000,000
0
@ /0,000,000
o
4
2,000,000
-8
v
6,000,000
-8
v
10,000,000
-8
1
*
.
.
I
-.4
t-H
t
71TTtt-1
L
‘.6
-8
0
-4
Angle
Figure
24-
mounted
M,
0.20.
The lift
ab we
4
8
of attack,
characteristics
the fuselage
/2
16
20
for ~ = 4“
a, &g
of i’he airplane model
and with the flaps
with the horizontal
o’eflecfeo! an, 30°;
~f 50°;
Iaij
1
.
101
NACA RM A9101
0
-.2
-.4
-. 6
0
.04
.08
./2
./6 ,20
Drug
Figure
25.-
8f, 50:
obove
Iul 0.20.
.28 .32 .36
for~=4°
coeffickm~ CD
The drag churacferistics
tail mounted
.24
of the uirplone
model with the horizontal
the fuse loge and with the flops deflecfed
#n,30°:
-
102
NACA FM A9101
.
.
—
,..&
--
‘... .*
.*
,.!.
.
..
:-.
—
.7
.6
.5
4
.3
.2
Pitching+wnmot
J
O
coefickw~
-J
?2
-3
-.4
~
of the oirplone model with the
26.- The pitching -mome@ characteristics
horizontal tail niounted ubove the fuseluge and with the flops deflected,
Figure
4,30;
4,50°;
M, 0.20.
-5
,
1
H
1
1
I
‘Q
.
.~~
-0--
~oxjmum ~s~
-!3--
Wake boundaries
---—
Horizontof
-80
tail
‘. \
—Horizontal
‘j
in the
e.sh?ndad wing-chwd
—
\
[email protected]
tad abova
##one
the
‘x,
•. \
‘\
1 \
I
fusebge
+(x)’
16-8
-8404812
Angle
of dock,
a,
4
O@
0
Angie
2Z - The variation
horizontal-tail
positkms
with Ongie of attack
with reference
4
c# dto~
I
I
8
L?
16
a, deg
(b) M, 0.50.
(a) M. 0.20.
Figure
I
of the
wake-boundary,
to the wing-chord
pfane
maxiwm-pressure-fess
d
a
=OO. Wing
and
locations
fuselage.
ond
R, 2,000,000.
-.
?-4
8
-8-404
Angle
of otfad,
12
B
0
/@71e of aftock,
8
u,
/2
deg
U, d~
16
-84
Angle
(~M,
(’C)IW, 070.
Figure
4
&280.
.
0
4
8f216~B
ofottock,
u“, deg
(e)tW, 0.85
v
27.-i2xv’inued.
.
.
,1’!’!
1’
.
,
E!
I\,
,
fuselage
\\
\\
\
\
\
1
\
A
\
E
-7
\
+
1’ ‘ \
\
\lz
\
“
I
~\
\
\
\
Y
\“
\
/
1
.1
\
\
\
r
\
H--H
—
‘8
An@
Figure27
.–
of
aftoc~
(f)
.44,0.90.
Conclu&d.
a,
d?p
0
-4
Angh
-=z?@-2
d
(d
4
ottook, u, &g
M ,0.95.
8
106
NACARM
A9101
.
80
60
Eli
\
~
R
2,000,000
#axijnum pressure /0ss
~
6,000,000
Maximum pressure loss
*108000,000
~
28000,000
Mt3ximumpressure /0ss
Wake boundaries
~
6,000,000
~/0,000,
000
Wake boundaries
Woke boundaries
.
.
.
‘8
-4
0
4
8
/2
16
20
Angle of ottack, U, o’eg
Figure 28. - The variation with ongle of uttock of the woke-boundary,
maximum-pressure-loss locations ond horizQnfal- tail positions with
reference to the wing-chord plane at u” O‘. Wing ond fuselage
with the flaps deflected
%n,30~ %t,50~ M, 0.20.
NACARM
107
A9101
.241
0
Wing ulone
1
.20
./6
./2
“
.
.
.08
.04
0
,
/0
Disfonce a..ove wing -chord planeat
Flgufe 29. - Wing woke
semitdl
areu.
~= O”, percenf ct
patterns
at the cenfroht
Q, 61 M, 0.85; R, 2,000,000.
.
.
of
108
NACA RM A9101
.
/,00
.98
.96
.94
.92
.90
.88
.86
.84
o
.3
*
FFFF
-6
‘4
-2
(o) Horizonfu/
0
P
4
6
8
10
toil mounted in the extended wing-chord plone.
/2
,
Figure 30.- The variation with ong/a of ottuok of the rotio of the dynumic pressure
at the Iuil to free -stream dynamic pressure. R, 2,000,000.
NACA RM A9101
log
a
.
.
/.00
/
.98
\
l\\
\
.96
\
.94
\\\
,
1
.92
.
.90
.
.88
.86
‘6
-4
-2
Angle
0
of
2
6
4
affoc~a,okg
(b} Horizon?oi foil mounted obove the fuseluge.
Figure 30.- Concluded.
.
8
10
/2
c
110
NACA RM A9101
.
.
.
.
Figure
3/ .- The variution
pressure
&
300;+,
ut the
50:
fail
with ungle of attack
to free -streum
of the rutio
dynumic
M, 0.20.
-QzEis2$
pressure.
of the dynamic
Flups
deflected.
NACARM
111
A9101
●
.90
●
.86
.82
.
.78
1
.74
.70
.66
.62
.
.58
.54
ti
.50 I
.
I
I
1
1
1
I1
1
1
1
1
m
llb&Ll_lll
I
I
I
1
I
I
I
1
I
I
I
I
I
I
I
Iu=o.50
/ -
.46
.
.42
.38
. . .
.34
.30
.26
.22
H+H+H+t-+t-tti
I
--
A4=0.20
.18
-8-6-4-20246
8
/0
Angle of ai’fuck, a, deg
(0) Horizontal
Figure
32.-
d attack.
tail mounted in the extended wing-chord phne.
The variation of Mach number at the toil with angle
112
NACA NM A9101
.
.98
A4=o.95
.94
.
\
M=O.92
~.o.go
.90 — “
MnO.85
.86
.82
M=O.80
k
.78
.74
-.
-.
.70
~.o.
TT-Kr”
70
I 1
.66
●
.62
✎✌
.52
M=O.50
.48
\
.44
.24
.20 ‘
\
M=o.zo
v
—
./6
-6
-4
-2
Angle
(b)
Figure
32.-
Horizonful
Concluded.
0
of
attock,
2
a,
toil mounted
4
6
8
o%g
above the fuseloge.
10
113
NAC!ARM A9101
/
/
.
I
L4-
I
-41
‘8
(0) Hori20nt01
Figure
33. -
angle
centroid
of
1.4’
I
-4
Angle
toil
I
of uffffc~
mounted
down wush at
semitoil
i
I
0
The voriation
of the
I
in’ the
80
*0.
‘
ml
I
I
I
8
4
I
/2
a, o’eg
extendeti
with angle
o position
wing-chord
of attack
of the
corresponding
oreu. R, 2,000,000.
plone.
effective
to the
114
NACA RM A9101
.
“
‘
‘
~
/
~----
----
.#--
-----
I
-;=O.sO
-
/
/‘
. /
M“O.20
#---.///- --I
~“ 0.70
I
‘
I
M“O.80
, ““’
#“
g
.
.
.
‘8
-4
0
Angle
figure
of attack,
toil mounted
(6)
Hort’zonful
33.
-Concluded.
4
a,
8
/2
f~e
fuseloge,
deg
ti&oVe
.
.
.. \.._..,
~z!E!3b
.
115
liACARM A9101
‘2 [
R
-2,000,000
----6,000,000
—+0,000,000
.
8
.—. ---A
4
0
4
(a)
‘8
-4
Horizontal
Angle of attack, e,deg
toil mounted In the extended
8
/2
wing-chord plane.
4
0
-4
-8
4
0
8
/2
Angle of attack, a, deg
(b) Horizontal
Figure 34?tfownwash
area.
Flops
The variation
of
oposition
de ffected.
tail mounted above the fuselage.
with angle of utfuck
corresponding
&,30e;
c?f,50”;
to
the
of the effective
cenfroldof
M90.Z0.
the
ongle of
semifuil
.. ..
.
,,
(o)
(2
./
Horkontol
fail
mounted
.3
.2
in the
.4
.5
number,
Moth
(b)
Figure
35.-
The variation
Wth
Horikon?oi
Mach
number
tail
of
mounted
the
.
extendd
.6
efficiency
.
@ne.
.7
.8
.9
Lo
M
obove
tail
wing-chord
fuselage.
factor.
*
NACA RM A9101
.
.
/.0
‘
.
R
—
2,000,000
-----6,000,000
— —/0,000,000
.9 ‘
v qyq
.8
.7
~a) Horizontal
fail mounfed in the exfended wing-chord plone.
.9
- / - -. -----
--- \
t-
*.M “
.8
v ‘Yq
‘
.e@H
—--.
-
:
/~
-–
.7
‘8
-4
.4
0
Angle of offack, a, deg
8
/2
lb) Horizontal fail mounfed above the fuselage.
Figure 36.-
The variation
Flaps de flecfed, {,30~+,
of fhe fail efficiency
50~M,0.20.
factor
with angle ofatfock.
-.
.-
8
~ s0.6
. -.
‘
—.
.
-
---
—
. -..
----
_
-
—
.
i
o,~
-
0.3
.
—_
-.
.
-.
--. _
.
.
_
_
:
-.
--
=
a
.
—
0.2
.
__
of ---– -:-.–--—— -----.-—-- ----------------....----- —----..---_ _...
_._. ___ .
~.-
0
8
.-
s
CL 0“6
---
.-
-. .
;;4
--
--
--
-
--
‘-
-
-- --
4
I--H-’”zr
1
1
of
J
“o
----
------
.4?
----
------
Mach
Figure
37.—
coefflclenfs.
The
varivtion
R, 2,00i9,00~
with
tail
Mach
it,
0?
---
.4
..3
[b) Horizontal
---
‘-*’-
WWUIJ
I uu
’-J
number
----
-_
uuuw
‘A-”’-
of the
-----
---
\
. ---
/
--
-.
—
-,
----
----
—--— ------
.7
.6
.5
number,
II
---
----
.9
.8
i.o
M
trtw
‘k-
angle
usaruyu.
‘--I---
=
J
of attack
of
the
airplane
model
for
various
Iif t
E
El
&
H
o
P
E!
.
Mach
(b)
Figure
38.-
?%e variation
coefficients.
Horizontal
tail
with Mach
R, 2, 000,000;
mounted
n:mber
~, O.
of
number,
above
the drag
M
the
fuselage,
coefficient
of the
airplane
model
fir
various
lift
--
kj
ii
8
Wlforizontol
tail
mounted
in the
extended
wing-chord
plane.
J
-----
y
G“
“ ;:?
.0
.
-------”
-----”
----”
‘----
“--=--
“----
-----
‘“--- “--
~
~o
___
r
0.4
‘-
-
,----
--
//.
--
0.3
Q
u
-----
- “—
‘--
—
~
~ ---
/
‘-
. .. . .
Y
“
Y,
.
‘..
,
\
3
$-.1
_
-.
_-
0.6
----
1
~
~
..
K
=&=–
-.2
o
.2
./
.3
Mach
(b)
Figure
39:
airpfune
t.,
.4
.3
The
variation
modet
for
Horizontal
tail
with
number
varibus
Moth
lift
mounted
coefficients.
r
.6
number,
,
of the
.7
,8
.9
Lo
M
above
the
fiselage.
pitching-moment
R, 2,000,00~
~, O:
coefficient
of jhe
.
.
$?4
II
-
b’
4s’
9
I
.2
----
-----
(a) Horizontal
Illil
_
-
-
Q,6
I
-----
-----
toil
mounted
-----
----
----
in the
----
I
I
------------
extended
I
I
---
wing-chord
I
—-/----
I
“-
1
1
plane.
.
/
CL*.6
.4
.2
----
0,
---
--—
----
----
----
____ ----
----
----
--
- -----
----
~
----
-----
J
\
o
o
.2
J
.3
.4
.5
.6
Mach
(b) Horizontal
Figure 40.-
The mriotion
corresponding
to
the
with
Mach
centroid
number
of the semitoil
number,
tail mounted
of
the
ores
.7
.9
lo
M
above
effective
for
.8
various
the fuselage,
angle
lift
of
downwosh
coefficients,
at
a
position
R, 2,000,000.
~
i’-
Iv
w
1.0
.8
!!.
.
HorizontoJ
toil
mounted
in the
extended
~
~
~
-
Horiz
-----
ontal
-----
toil moun ted above
.- —- -----------
I
I
I
-
-----
1
plane
~
~
~
wfng-chord
~
~
----the fusela
-----
go
----
----
.-:--
-a--
I
I
[
I
0
0
I
1
I
.3
.t?
J
I
I
I
.4
.5
Mach
Figure
4/ .–
ifownwosh
The
with
variation
angle
with
of
Mach
attuck.
number
I
of
I
.6
number,
/he
I
rate
I
.7
.8
‘--
I
1
I
I
I
.9
1
I
10
M
of
change
of
the
effective
angle
~
of
E$!
g
1+
R, 2,000,000
~
,.
,
,
NACARM
123
A9101
.8
,.
Qti
\
/“
.6
4=-4;
__ ----------- --—---------------3 ‘-- ---/ /
.4
,/
-2°
s
.; .2
Q
“w
k
:0
Q
~
~ -.2
-/ 0
-
0°
‘-
-
\
/ ‘ . \
/
4
/
u
6 /
~
‘-
K
‘--
‘-
-
\
1,-)
-.4
-.6
~a) Horlzonfal tailmounted
.
intheexfended
wing-chord p/a;e.
.8
.
iL-40
t --3&
-2°
.6
●
(y .4
*
~.
-=
q
~----___---- --------- ---- ----- --- _&,
I
‘k
.
‘
-I
I
I
0°
1
1
-
$ ‘l-t-t-t’:
0
I
~
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I
1
1
1
1
-
1
I
,
—
/
,
“
1
b-l
—-
l— ~.
4°~
I
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‘--
2L
,
r \
--
—
-
--.
I
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.
Y
%
/
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,I
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=?3=
.6
0
J
.2
.3
.4
.5
Much number,
@}
Horizontoltoil
mounfed
.6
.7
.8
.9
lo
Af
a&ovefhe
fuseioge.
Figure 42.- The vorjation of fhe /jft Coefficient for Cm=O with Mach number.
R, gooqooo.
.-
-.
.8
\
\
\
\
\
\
rl=4
~,.
b,
.6
. 4
II
.2
0
.5
--H--H-W
,6
.7
.8
Mwh
Figure
wing
43. - The lift
loading
an altitude
,.
of
of
coefficient
100
pounds
10,000
number
requt’red
per
.9
10
~ M
for
square
an airplane
foot
in fli&f
with
a
at
feet.
.
.
.
,
o
z!
b
%
+
-2
*
$
~
-4
(o) Horizonto/
toil mounted
in the ~Ambd
wi~-chord
plane.
Z
4
k
%
*
a
32
\
o
*
*O
v
-2
-4
0
.1
.2
.3
,5
.4
Abchnmhw,
(b)
~igtue 44. - Variation
h
flight
Ot
laOOO
with
Hor’izontol
Mach
feet.
number
Wing
100ding
the
100
.8
.7
10
.9
M
toll mounted
of
.6
angle
rends
above
the fuseioge.
of stobillzer
Per
Wu~e
seftlng
f~.
to balance
Genter
of 9fovi!Y
an
olrpione
ot
O“25c’=
.-
.8
c?
.4
-.4
th)M,
(NM,
a2t2
{i#M.
0.50.
0.80.
.8
.6
-.4-
8
-4
0
4
-8-404
-8-404
Elevator
fb)hf,
0.85.
Figure 45. - 7i?e lift characteristics
elemnbr.
.
.
(f)hf,
deflecfion,~,
0.90.
of the imlotd
R, ~O@700.
.,
-8404
deg
@M,
0.92.
toil with a 20-percent-area
6)M,0.94.
cmsfant-chcwd
,.
.
‘.
.
,
---
f)=o
.. ~
L
~-
.
/“
#
/“
/
. .
-
/
/
/
G-z
.6;
c=== ,~
,
—
-.
/7.84
(
/
I
(
n=3
{o) Horizontal
mounted
in
the
extended
/).
I
~
?!
w
toil
-2
wing-c40rd
I
1
1
I
I
I ..-++-
plane.
1
I
*-
{, O:
1
Y
kt=l
I
I
-4
-6
0
.1
.2
,6
.3
Ah
(b) Horizontal
Figure
46.-
at 10,000
Voriotion
tith
Moth
feet. Wing ioading
number
ms&r,
toil mounted
of
JOO pounds
the e/evator
per
squore
.7
,8
1,0
.9
M
ofwve
the fuselage.
deflection
foot. Center
to bolance
of gravity
+,2:
an airplane
at 0.25
c!
in f/ight
—.
I
u
“ill
03
.L
4
.4ngle of stabilizer
Figure
47.-
The angle
the horizontal
ofstobifizer
tail mounted
0
-4
0
sefting,~,
setting
in tie
.04
deg
for
extended
.08
.
GO
and
wi~-chord
iWg
Me
.12
coefficient,
corresponding
plane
.16
and the
.24
.20
CD
drag
flaps
coefficient
deflected.
for
~,
30°;
modef
the
~,
50°;
witi
&l-l
M, O.20.
0
1-
..
.
+
.
,
v
,
.
II
-2
0
Lift
Figure
in
I
48.
the
–
The
extended
lift-drug
wing-chord
rqtio
.6
.4
.8?
coefficient,
of
the,
pfune.
.8
/.4
CL
airplane
8n, 30°;
model
with
the
horizontal
~, 50?j (%, O ; M, 0.20.
tail
mounted
130
l?ACARM A910~
,
?
-+
160
-1-L2rL
200
240
Gliding speed, mph
.
280
Figure 49.- Variation with gliding speed of the sinking
qveed of on airphne in power-off
flight at sea
level. Wing loading 100pounds per squure foot.
.
Center of gravity at O.~Sc!
~
1
b
“,
9
._
.
4
NACA-Langley
-12-8-49-
S’75