Investigation of effect of sideslip on lateral stability characteristics

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Title:
Investigation of effect of sideslip on lateral stability characteristics
Alternate Title:
NACA wartime reports
Physical Description:
15, 18 p. : ill. ; 28 cm.
Language:
English
Creator:
Fehlner, Leo F
MacLachlan, Robert
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:
Airplanes -- Fuselage -- Testing   ( lcsh )
Airplanes -- Wings -- Testing   ( lcsh )
Aerodynamics -- Research   ( lcsh )
Genre:
federal government publication   ( marcgt )
bibliography   ( marcgt )
technical report   ( marcgt )
non-fiction   ( marcgt )

Notes

Summary:
Summary: The results of tests of a circular fuselage with various combinations of tail lengths and vertical tail surfaces with and without the horizontal tail surface in the 6- by 6-foot test section of the NACA stability tunnel are reported in the form of diagrams of variation of coefficients of lateral force and yawing moment with angle of yaw and angle of attack. The results of these tests indicated that the change in the unstable yawing moment of the fuselage alone due to increased tail length did not appreciably affect the yawing moment of a fuselage and vertical-tail combination. The addition of a horizontal tail increased the efficiency of the vertical tail in normal-flight attitudes and in the region of negative angles of attack. Existing methods of computing tail effectiveness gave results within ±7 percent of the measured values for the cases computed.
Bibliography:
Includes bibliographic references (p. 12).
Statement of Responsibility:
by Leo F. Fehlner and Robert MacLachlan.
General Note:
"Report no. L-12."
General Note:
"Originally issued May 1944 as Advance Restricted Report L4E25."
General Note:
"Report date May 1944."
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."

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University of Florida
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All applicable rights reserved by the source institution and holding location.
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aleph - 003619649
oclc - 71360411
sobekcm - AA00006287_00001
System ID:
AA00006287:00001

Full Text

(rmACrL-1y- V




NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS





WAlRTIM E I RE PORT
ORIGINALLY ISSUED
May 1944 as
Advance Restricted Report L4E25

INVESTIGATION OF EFFECT OF SIDESLIP ON
LATERAL STABILITY CHARACTERISTICS
I CIRCULAR FUSELAGE WITH VARIATIONS IN
VERTICAL-TAIL AREA AND TAIL LINGTH
4 WITH AND WITHOUT HORIZONTAL TAIL SURFACE
By Leo F. Fehlner and Robert MacLachlan

Langley Memorial Aeronautical Laboratory
LanglEy Field, Va.









WASHINGTON

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 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.

D DOCUMENTS DEPARTMENT
V. L-12






































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/investigationofe001a




-7


NACA ARR No. L4E25 RESTRICTED

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS


ADVANCE RESTRICTED REPORT


INVESTIGATION OF EFFECT OF SIDESLIP ON

LATEaAL STABILITY CHARACTERISTICS

I CIRCULAR FUSELAGE WITH VARIATIONS IN

VERTICAL-TAIL AREA AIID TaIL LENGTH

WITH AND WITHOUT HORIZO':TAL TAIL SURFACE

Ey Leo F. Fehlner and Robert MacLachlan


SUMMARY


The results of tests of a circular fuselage with
various combinations of tail lnryths and vertical tall
surfaces with and without the horizontal tail surface
in the 6- by E-foot test section of the NACA stability
tunnel are reported in the form of diagrams of vari-
ation of coefficients of lateral force and yawing
moment with angle of yaw and angle of attack.

The results of these tests indicated that the
change in the unstable yawing mor..ent of the fuselage
alone due to increased tail length did not appreciably
affect the yawing moment of a fuselage and vertical-
tail combination. The addition of a horizontal tail
increased the efficiency of the vertical tail in
normal-flight attitudes and in the region of negative
angles of attack. Existing methods of computing tail
effectiveness gave results within 7 percent of the
measured values for the cases computed.


INTRODUCTION


Desirable qualities for th3 lateral stability and
control characteristics of an airplane are dependent
on the set of stability derivatives peculiar to the
airplane. The stability derivatives can be changed
by changes in airplane parameters, such as vertical-
tail area, horizontal-tail area, end tail length.


RESTRICTED








iU.CA ARR No. L4S25


Extensive tests to determine the changes in stability
derivatives effected by uniform changes in airplane
parameters have been made with a model geometrically
similar to the model used in the present investigation.
Included in these tests were the effects of cowlings,
of wing positions, and of the presence of a vertical
tail (references 1 and 2). Reference 1 is mainly con-
cerned with lift and drag characteristics, whereas
reference 2 deals with the effects of yaw on the
lateral stability characteristics of a rectangular
wing with a circular fuselage and vertical tail.

The present investigation, in which the model
was tested without wings, is an attempt to determine
experimentally the basic changes in stability deriva-
tives caused by uniform changes in vertical-tail area
and tail length and by the presence of a horizontal
tail. Because a geometrically similar model has been
tested in the LMAL 7- by 10-foot tunnel (reference 3),
the data may be used for correlating the results in
the two wind tunnels.

The tests were made in the TACA stability tunnel
and included an angle-of-attack range from -100 to 200
and an angle-of-yaw range from 120 to -300 with various
combinations of three fuselages of different lengths
and five vertical-tail areas with and without a
horizontal tail surface. Two combinations of the model
parts used in the present tests are geometrically
similar to the model used in the LMAL 7- by 10-foot
tunnel for the tests of references 2 and 3.


APPARATUS AND MODEL


The tests were made in the UhCA stability tunnel
6- by 6-foot closed-throat test section with the
regular six-component balance.

The principal dimensions and the arrangement of
the parts of the model used in the investigation are
shown in figure 1. All the model parts are made of
laminated mahogany. Figure 2 shows the model
unassembled, and figure 3 shows the model mounted
on the model support. The horizontal strut supporting
the model does not rotate in pitch with the model.
The vertical struts rotate in yaw with the model and
remain alined with the relative wind.








NACA ARR No. L4S25


Th3 fusalace is cf circul.sr cross section. Its
length can o c: ,-:;d b7 i : si.- of thr-e i:-= changeable
tail cones. : ,* :.r cf the-; .-. "L c nes is
attached, th: f fi~ i -=.: tr.icall .; ir iar to the
circu.l'r fts la.-. ,d c r : n re.1 rn" 1 and to the
mo-lel u.3dd for t:' t.e-:I : _r-.'red in rere i ne 2. The
co.': :a tces for the m:.-n, I axTi d long tail con2 s were
obtained by extending Lhe acscissa for the length of
the short tail cone according to the formula

S= X + X (c 1) sin a X
where

Xo abscissa for original length

a length of portion to be distorted

c ratio of original length of portion to be distorted
to final length of portion distorted

X1 abscissa for final length of distorted portion

The ordinates corresponding to X1 are taken as those
corresponding to Xo from which X1 was computed. The
tail Isngths, the lengths of the three fuselages and
tail cones, and the ratios of the tail lengths to the
48-inch span of the proposed wing are given in table 1.

Five geometrically similar vertical tail surfaces
were made to conform to the NACA 0009 section. In plan
form they are representative of the vertical tail
surfaces used on the average airplane. The geometric
aspect ratio of each vertical tail is 2.15. The
horizontal tail surface was made to conform to the
NACA 0009 section. Its geometric aspect ratio is 3.99.
The numbers by which the tail surfaces are designated,
their areas, and the ratios of thest areas to a proposed
rectangular wing area of 361 square inches are given in
table 2.


TESTS


The model combinations tested are given in table 3.

Angle-of-attack tests for each model combination
were made over a range from -100 to 200 at angles of yaw







T ACA ARR No. L4E25


of -50, 0, and 50. Angle-of-yaw tests for each model
combination were made over a range from 120 to -300 at
angles of attack of -100, 00, 100, and 200.

The dynamic pressure for the tests was 65 pounds
per square foot, which corresponds to a velocity of
about 160 miles per hour. The Reynolds number based
on an 8-inch wing chord was about 888,000.


RESULTS


I.1- results are presented as standard nondimensional
coefficients based on the dimensions of a rectangular
wing proposed for the model. The following symbols are
used herein and the senses are defined relative to a
person within the airplane facing the direction of motion:

Cy lateral-force coefficient (Y/qS,)

Cn yawing-moment coefficient (N/qSwb)

Y lateral force (positive to right)

N yawing moment (positive when right wing tip tends
to move rearward)

q dynamic pressure ( pV2)

p air density

V tunnel-air velocity

Cy =Cy
Cy n

6Cn

angle of yaw, degrees (positive when right wing
tip has moved rearward)

a angle of attack, degrees (positive when tail has
been depressed)

L tail length

b wing span (48 in.)


Sf vertical-tail area









HACA ARR ITo. L4E25


Sg horizontal-tail area

S, wing area (-61 sq in.)

Af aspect ratio of vertical tail surface

Figure 4 shows the system of axes used in the
measurement of forces, moriients, and angles. Tns axes
are fixed in the model fo~' all changes in angle of yaw.
Ecr changes in angle of attack, the X-axis remains
in the plane in which it was located at a = 0. The
axes intersect the rmoe l at the assumed center of
gravity, which is 10.40 inches behind the nose.

The lateral-stability derivatives are computed,
for the range of an;le of attac-:, ir,:m measurements
of lateral force and yawing moment at angles of yaw
of 50; the variation of the forces and moments with
angle of yaw is assumed ti te linear over the 50 range
of angle of yaw.

Angle-of-yaw tests were made to check the linearity
of the curves of Cy and Cn against angle of yaw in
the -5 angle-of-yaw range. The slope of these curves
shows that the variation of the forces and moments
within the angle-of-yaw range of 50 is linear except
at high angles of attack. The measured slopes of these
curves are plotted with tailed symbols in the figures.

The measurements of lateral-force coefficient Cy
are considered accurate to 0.0012 and of yawing-moment
coefficient Cn to 0.0005. The ang.le-of-yaw measure-
ments are accurate to about 0.050, and the angle of
attack is accurate to about 0.10.

A model geometrically similar to the IACA stability
tunnel model was tested in the LMhL 7- by 10-foot tunnel
and the results of the tests were reported in refer-
ence 3. The model consisted of the short fuselage,
vertical tail surface 4, and the horizontal tail surface
and was tested in the LMAL 7- by 10-foot tunnel at a
dynamic pressure of 16.37 pounds per square foot, which
corresponds to a velocity of 80 miles per hour. The
Reynolds number based on a 10-inch chord was 619,000 and
the turbulence factor was 1.6. The model tested in the
NACA stability tunnel is eight-tenths the size of the
model tested in the LMAL 7- by 10-foot tunnel and was








rT.AC:. ARR i'o. L4E25


tested at a dynamic pressure of 65 pc'jnds per square
foot corresnondinr! to a velocity of 160 miles per hour
and a Reynolds runmber of E-3,000, The turbulence of
the air stream in the NACA stability tunnel is not
known but is believed to be lower than that of the
LMAL 7- by 10-foot tunnel. Variation of CyI and Cn,
with a for the similar rncesls are shown in figure 5.
Values of Cy and Cn agree well for the two sets of
data. Tne maximum discrepancies occur at high angles
of attack in the re-ion of the stall.

In order to check the data obtained in the
ITACA stability tunnel, a temporary one-co:ironent spring
balance was installed to measure the yawing moment due
to sideslip. The model support consisted of a cylindrical
rod fixed perpendicular to the top of the tunnel by a
tripodal wire stay. The model was supported in the same
position in the tunnel as on the regular tunnel balance
except that it was inverted. Such an arrangement was
expected to give altogether different interference
effects from the regular support. Filur. 6 show': the
variation of Cn with thus obtained for a typical
case at an ar.gle of attack of 00 and ccmc-ared with
similar data for the model on the reu-:lar support in
the NACA stability tunnel and in the LMAL 7- bI 10-foot
tunnel. The two sets of data obtained in the M-.CA sta-
bility tunnel check each cther. The data obtained from
the LMAL 7- by 10-foot tunnel check the slope from the
NACA stability tunnel within 8 percent. This difference
in slore is the same as the difference shown in fig-
ure 5 for C at an angle of attack ef 00. The source
of the discrepancy is not obvious from the data tut may
be the differences in the deflection of the models,
angularities of the air stream, or model-size to jet-
size ratios.

T7'E results are presented in the form of curves
that show the effects of chan-res in fuselage length for
fuselage alone in figures 7 and 8; of ch-eaires in
vertical-tail area, figures 9 and 10; of changes in
tail le- th, figures 11 to 13; of addinL the horizontal
tail surface, figures 14 to 16; and of charnes in tail
length and vertical-tail area with constant tail volume,
figures 17 and 18. The data lctted in the various
figures and the model combinations used in obtaining
each plot are summarized in table 4.









ICACA ARR No. L4E25


DISCUS3IC'

Effect of Changes in Fuselage Length for Fuselage Alone


The effect of changes in fuselage length on the
stability of the fuselage alone is shown in figures 7
and 8. The derivative Cy,, was increased by an
increment of 0.0007 by increasing the value of i/b
from 0.418 to 0.618. The absolute magnitude of this
increment is small compared with the magnitude of
derivatives for the fuselage with the vertical tail
surfaces tested. The derivative Cn is very slightly
changed by the change in fuselage length relative to the
magnitude of the derivatives for the complete model.
Although theoretical analysis indicates that the unstable
yawing moment of the fuselage alone varies with fineness
ratio and volume, this variation has not been detected
herein because the magnitude of the variation is of the
same order as that of the experimental accuracy.


Effect of Changes in Vertical-Tail Area

The effect of changes in vertical-tail area (hori-
zontal tail on) is shown in figures 9 and 10. At an
angle of yaw of -100 and at an angle of attack of 0,
changing Sf/Sw from a value of 0.0659 to 0.0974
increased Cy by an increment of 0.019 and Cn by an
increment of O.C097. Throughout the angle-of-attack
range, the same change in Sf/Sw increased Cy by
an increment of about 0.0014 and Cn by an increment
of about 0.00098.

The values of Cy and Cn decrease with angle
of attack; the decrease for a change in angle of attack
from -5 to 5 is 0.00043 for Cn with vertical tail
surface 2 and 0.00048 with vertical tail surface 4. The
decrease in Cy, for the same decrease in angle of
attack and for either vertical-tail area is 0.0012.









Ir.CA ARR No. L4E25


Effect of Changes in Tail Length

Tl-. effect of changes in tail length is shown in
figures 11 to 15 for the model with the horizontal tail
surface and vertical tail surface 4. Th= change in Cy
due to changing L/b from 0.418 to 0.618 is small and
probably negligible for cases in which the lateral force
is largely the contribution of the vertical tail surface.
The effect on Cy as shown in fi'ure 12 appears
-f
inconsistent but is small and therefore probably
negligible.

The yawing moment due to sideslip increases with
tail length. This increase in Cn increases with '
up to shortly after the stall of the vertical tail sur-
face. At values of & beyond the stall, Cn is
increased about 0.01 by an increase in tail length
of 0.1.

Changing the value of L/b frorr 0.418 to 0,518
causes an increase in Cn of approximately 0.0007
throughout the angle-of-attack range. An increase in
i/b, however, from 0.518 to 0.618 causes increases
of 0.00059 and 0.00046 in Cn, at angles of attack
of -50 and 50, respectively. Increasing the angle of
attack decreases Cn For the short, medium, and
long tail lengths, the decrease in On is 0.00015,
0. _0027, and 0.00050, respectively, for an increase in
angle of attack from -50 to 50.

The effect on Cn and Cy, of changing the
vertical-tail area and tail length is shown in
fi I T1. Ti.e model, in this case, is at an angle of
attack of 20 and is equipped with the horizontal tail
surface. Increases in vertical-tail area produce
regular increases in both Cy and C, Increases
in tail length produce regular increases in Cn
except for the extremely small values of Sf/Sw for
which the directional instability is of the same
order of n:ranitude as for the fuselage alone. For
all practical values of Sf/S,, therefore, increasing









I'ACA i.RR No. L4E25


tuil length increases the directional stability as
measured by ,C .


Effect of Horizontal Tail Surface

The removal of the horizontal tail surface decreases
the efficiency of the vertical tail surface in all
attitudes except at large angles of attack. (Sea figs. 14
to 16.) Fcr the long fuselage, at -50 angle of attack,
the value of Cyl, is decreased by an increment of 0.001
by removing the horizontal tail surface whereas, at
50 angle of attack, Cy, is not decreased. The effect
on Cy,, in reneral, is the same magnitude for the
short-fuselage and vertical-tail onr.binaticn. For the
long fuselage, figure 15 shows a large effect on Cr

that varies from a decrease of 0.00090 at an angle of
attack of -50 to a decrease of 0.0'.JOJ9 at an angle of
attack of 5. The corresponding decreases for the
short-fuselage and vertical-tail combination are 0.00054
and 0.00020. (See fig.16.) By removing the horizontal
tail surface, the efficiency of the vertical tail
surface is therefore decreased by an amount that varies
with angle of attack. The decrease is slightly greater
for the short-fuselage and vertical-tail combination
than for the long-fuselage and vertical-tail combination.


Effect of Changes with Constant Tail Volume

The effect of changes in tail length an! vertical-
tail area with tail volume held constant based on the
dimensions given for the model is shown in figures 17
and 18. The derivative Cy increases as the vertical-
tail area increases and as the tail length decreases.
The value of the derivative Cn, theoretically should
not vary with changes in tail length and vertical-tail
area if the tail volume is held constant. The variation
of Cn0, measured experimentally, is small over the
range of angle of attack from 40 to 120 but is appreciable
at negative and at high positive angles of attack.








HIACL A.RR No. L4E25


Ccmcarison with Theoretical Variations

The experimental results have been compared with
theory by means of accepted simple calculations that
involve a minimum of anticipation for the desired results.
The first of these calculations can be made from the
expression of the variation of lateral force with side-
slip, which can be written as


CYtotal = *fuselage + (CY)f (1)

where Cy selae is the experimental value obtained
\ /fuselage
in this investigation,

Sf KAf
(C) = 2T Sw Af + 2K (2)

and f denotes vertical tail surface. The constant K
is given in reference 4 as 0.875 for an elliptical span-
wise loading. If the spanwise loading of the vertical
tail surface is assumed to be elliptical for the purposes
of analytical treatment and if the model dimensions are
used as previously given, values of Cy for vertical
tail surfaces 2 and 4 are 0.00348 and 0.00515, respec-
tively, according to equation (2). The corresponding
experimental values computed from the data according to
equation (1) at 00 angle of attack and with a horizontal
tail on the lone fuselage are 0.0038 and 0.0054. The
theoretical relation then underestimates the value
of Cy by 9 percent for vertical tail surface 2 and
5 percent for vertical tail surface 4. Similarly,
Cn may be written

( = ( ) (3)

Theoretical values of Cn for vertical tail surfaces 2
and 4 on the long fuselage are -0.00215 and -0.00317,
respectively. The corresponding experimental values for
the model with the horizontal tail surface are -0.00226
iid -0.D.-0?. T1h theoretical relation then under-
estimates the value of Cn for combinations with the
'1'








NACA ARR No. L4E25


long fuselage by 5 percent for vertical tail surface 2
and 4 percent for vertical tail surface 4.

The theoretical values of Cn, for vertical tail
surface 4 in combination with the medium and short
fuselages are -0.00267 and -0.00215, respectively. The
corresponding values of -0.00277 and -0.00207 were
obtained experimentally for the model with the
horizontal tail surface. The theoretical relation then
underestimates Cn, for the medium tail length and
vertical tail surface 4 by 4 percent and overestimates Cn
for the short tail length and vertical tail surface 4 by
4 percent. If the value of CyI commuted according to
equation (2) Is increased ty 2 percent, the resulting
values of Cy, and Cn,. are v'ithin 7 percent for the
cases analyzed. This 2-oerc!nt increase in Cy| may
account for the influences of the horizontal tail surface,
the influence c.f the fuselage, or any peculiarities of
flow. The resulti-ng discrepancies, which amount to
7 percent, are slightly less than twice the limits of
discrepancy shown previously between the data from the
NACA stability tunnel and the LW.L 7- by 10-foot tunnel.


CO:;CLUS IO3S


The results of tests in the NACA stability tunnel
of a circular fuselage with various combinations of tail
lengths and vertical tail surfaces with and without the
horizontal tail surfaces indicated the following
conclusions:

1. The effect of the change in the unstable yawing
moment of the fuselage alone due to increased tail length
on the variation cf yawing moment with sideslip was
negligible.

2. At an angle of attack of 20, the vertical tail
surface in the presence of the horizontal tail surface
produced values of lateral-force derivative Cy and
yawing-monent derivative Cn. that were within 7 per-
cent of the estimated values.








FACA ARR HTo. L4E25


3. The addition of the horizontal tail surface
increased the efl'ciency of the vertical tail surface.
The increase in Cy varied from 0.001 at an angle of
attack of -50 to 0 at an angle of attack of 50; and the
increase in C varied from 0.00090 at an angle of
nf*
attack of -5 to 0.00039 at an angle of attack of .


Langley Memorial Aeronautical Laboratory
NItional Advisory Committee for Aeronautics
Langley Field, Va.





1. Jacobs, Eastman N., and Ward, Kenneth E.: Inter-
ference of Wing and Fuselage from Tests of
209 Combinations in the N.A.C.A. Variable-Density
Tunnel. NACA Rep. I1o. 540, 1935.

2. Banber, IT. J., and House, R. 0.: Wind-Tunrnel
Investigation of Effect of Yaw on Lateral-Stability
Characteristics. II Rectangular N.A.C.A.
23012 Wi-' with a Circular Fuselage and a Fin.
ITACA TN No. 730, 1939.

3. Wallace, Arthur R., and Turner, Thonras R.: Wind-
Tunnel Investigation of Effect of Yaw on Lateral-
Stability Characteristics. V S' ---trically
Tapered Wing with a Circular Fuselage EHvirng a
Horizontal and a Vertical Tail. TACA ARR
1o., .P23, 1943.

4. Higfl ns, George J.: The Prediction of Airfoil
Characteristics. NACA Rep. I:o. 312, 1929.







IT'.C0 ARR No. L4E25



''S TAB-L 1 SIOS

FUSELACE DIr.Mu3SIOUS


Fuselace rail-c~rne fail lenith, Tail len-th
(int .) (n .) (1)

Short 32.25 9.c5 20. 7 0.418

Mediun 37.C 14. .5 24. 7 .518

Long 41.85 1..45 2J.67 .616

Tail Isngth measured from center oif rLvvity, assi-uned to
te 10.40 i.n. behind .r..se of tnr nod.-l, t3 hin.e line
of tail surfaces.
I..iTIJ;kL ADV'ISOHY
C30,; ITT FOR AERONAUTICSS



TAELE 2


AREAS CF VERTICAL ha!D HORIZ,-'TL T.IL 3SJ1RChS

Tarsrfae Desgnaton Tail area Ta1l area
(sq in.) 'ing ar a

Vertical 1 10.83 0.0300

Do----- 2 2 .78 .C653

Do----- Z 28.Z7 .0786

Do----- 4 35.16 .0974

Do----- 5 46.20 .1278

Horizontal ----------- 64.21 .178











NACA AER I!o. L4'C25


,-4C



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,c a,-, r- C -- ~ L 1 1 L1 r./
I--





C. .. .1 C 0 0 0C 0 0 0 0 0 0 W
- ) %a LO I L l T) I' a i i I

.. I I I I I I I
I I I I I I I I I I


ID I I I I I I I I I "
., : : I I I I i I I I
4- ) I I I I I I I I

r-4 .-4 C 0 0 C 0 C' 0' 0 0 i 0"
r > t. 1 I 1 Z 0 1 1 1 1 -







-- I'I 'IO II '
S, I I r I I I IC






4t rI I I lI I rI IC C'


















L4, E
i- -
0







r .-J II I I0
0 0)I I I I I I I I I
." I I I I I I I I
0 -i 0 '0 0 C 0 0 0 0 0 0
--l i C I I I I I I I
SI I I I I I I I I


'D
1 I I -i-J I I I I 1 3
1-I r I ;:Q I -, I I I I I rt
'D 0 0 C 0 0 0 0 0 0 0
r :- 3 0 T .C 'U Ti 'U *TI V3 -
c i I ,4 I CO I I I I I ...
I I I I I I I I


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I.'C. A,RR No. L4E25





FFESE!TrTTCt" OF RESULTS


4 iod =1
Plot iod. 1 Picure
c :. J ir.at or

C a:nd C against a c9

Cn, against 9 6

Cy and Cn against 1 iad 3 7

Cv and rC, eagint a 1 ard ,

CY and Cr a.-ainst V 1?2 an 21 9

C and Cn against a 19 and 21 10

Cy and Cn against 9,,15, and 21 11
C 'and C, arsinst a 9,15, and 21 12
Sr
C and Cn., t

Sand C- at a = 2c L re 2" 12

Cy and Cn against 1' 4 and 21 14

Cy- and Cn against a 4 and 21 15

n
Cy and Cn against a 2 and 9 1i

Cy and Cn 4ainst ,14, na Q19 17

Cy and Cn against a 3,14, and 19 16


aAlso shown in this fi-ure are results from L :AL 7- by
10-foct tunnel fcr model with dl-.:escisns .eone-trically
simillar to .- ddl comtination 9 tested in U;ACH stability
t unnrl 1.


-.T IONAL AL1'T30I(Y
CO -'-T-- FO0 AERONAUTICS






NACA ARR No. L4E25


1. 15 0'R


NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS.


I.05JR H6.098


1.46 7 '"
0.711 "R -. 4.12


0.991 '


J 1.280 R 746'-


Lij IT784- \

T
---.-3.384


Figure I .- The circular fuselage, vertical and horizontal tails,

and tail cones with the principal dimensions for assembly.


Fig. I


1 Z80 R
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NACA ARR No. L4E25 Fig. 2







































o

E.
>

00




4..
Zu

22

I-I--
z j,
Z























































































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I







NACA ARR No. L4E25


NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS.


Figure 3.- Model mounted on model


Fig. 3


support.










NACA ARR No. L4E25 Fig. 4







>>
+-





C m
o o oZ











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Fig. 5 NACA ARR No. L4E25




--


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8 l: c<

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r- r-
--- --4'l -










I gs II


00 0







NACA ARR No. L4E25


NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS.


-30 -20 -/0 0 /0
Angle of yaw, t, dey


Figure 6.- Comparison of data obtained with NACA
stability tunnel regular balance, NACA stability
tunnel spring balance, and LMAL 7- by 10-foot
tunnel balance for rate of change of yawing-
moment coefficient with angle of yaw. Horizontal
tail surface on; 0- 0.118; vertical tail surface 4,
b
Sf = 0.0974; a = 0.
Sw


c
(a

+-

U
0


U

U,
E






3
r-t


.0/




0


Fig. 6






NACA ARR No. L4E25


%.



o. .

1-



.r


I2-





0? - -


o 0.418
+- .618


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,1ZZ


JILLEILI


-30


-20


-/0
Angle of


NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS.
2 0 20


yaw, p deq


Figure 7.- Effect of changing fuselage length on rate of change of
lateral-force and yawing-moment coefficients with angle of yaw.
No horizontal or vertical tail surfaces; a = 00.


r


F.ig. 7


I








NACA ARR No. L4E25


-- -- --- -- -










II



---<---Jo-
--- -- .) -- -- ---l



--- -- ) --- -- -- (C


fA3


Fig. 8


0
cUj



--J
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4r
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NACA ARR No. L4E25


q)


x 2- --








Vertical tail
--- -surface
0 0.0659 Z
+ _9Z4 4--








0
-o 0I0659 -
__ O 7 ._ _

>-









-W2I


-30


I NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS.


-/0 0 fI
Anyleofyaw U ,dea


Figure 9.- Effect of changing vertical-tail area on
rate of change of lateral-force and yawing-moment
coefficients with angle of yaw. Horizontal tail
surface on; = 0.618; a = 00.
b


Fig. 9







NACA ARR No. L4E25 Fig. 10



I-,






-- --- )--- g



04




i



o w t
o a

--- ;>---* ---- --c -5
H 4

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-4 tI



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--- 4 -
-0u 0


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Id
43
S"H









0









^0







NACA ARR No. L4E25


9 -10
Angle of yaw,


Fig. 11







O
(j
-^






a
a.,



a



4-

U
u










qj




o










C
0







4-


S,deg-


Figure 11.- Effect of changing tail length on rate of
change of lateral-force and yawing-moment coefficients
with angle of yaw. Horizontal tail surface on; verti-
cal tail surface 4, = 0.0974; a = 00
SW


NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS.


-30







NACA ARR No. L4E25 Fig. 12






LO



< I I
I--




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Z o
Wo
z4 0%
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O N
4 0 0





CO -r






NACA ARR No. L4E25


.0/



0



0
O



CAW -.002



-.004


.08

Vsw


NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS


Figure 13.- Effect of changing tail length on rate of
change of stability derivatives Oy* and Cn with
vertical-tail area. Horizontal tail surface on;
a = 20.


.04


.12


.16


Fig. 13







NACA ARR No.


L4E25


*O-- __ _


0---- 4- -





o o

.04
-+ .178











_---


-30


-zE


-10


NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS.


Fig. 14


Angle ofyaw, f/ deg

Figure 14.- Effect of horizontal tall surface on rate of
change of lateral-force and yawing-moment coefficients
lwth angle of yaw. A = 0.618; vertical tail surface 4,
b
f = 0.0974; a = 00.
Sw


e ,'








'ig. 15 NACA ARR No. L4E25


u

oo
00
TX CU






40
'\
Z 1



COd

I49

S 0 0



a.0
I '-
CIS 4- If
0 1





S I O 0 0 0o
,,_.v_' 1 ,, S .. 4










,, 0







Le as ,o
0\ f



0|
4O 0

0 '4

3 >



Erz




+ 0
0
0w








NACA ARR No. L4E25 Fig. 16



u
L>
>- I
Jr --- 44-
0o O




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I.d













0o ON
oc*

S o ;
0 )



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aa

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--- --- < -- --- -- -- cO g o c
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0 6) 0 0 o











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So





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1< 0 > B
I4- .-4 0






0 90





SII IE






O j o O
O O






NACA ARR No. L4F25


A- -------




O2--------------^^

Vertical tail
2/b ,J.5', surface
o 0.418 0.0974 4
x .518 .0786 3
+ .618 .0659 2














NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS
__


-30


-20


-10


Angle of yaw 9/, dey

Figure 17.- Effect of different combinations having
constant tail volume on rate of change of lateral-
force and yawing-moment coefficients with angle
of yaw. Horizontal tail surface on; a = o0.


Fig. 17








NACA ARR No. L4E25 Fig. 18




-i-




L.

O.bO
I-
o o
I-I-~





od

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tE
2 I X 4


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t





UNIVERSITY OF FLORIDA


3 1262 08104 976 8


.JiNi'ERSITY ) FLOOR IDA
DOCUMENTS DEPARTMENT
120 MARSTON SCIENCE LIBRARY
PO. BOX 117011
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