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:
18, 40 p. : ill. ; 28 cm.
Language:
English
Creator:
Hollingworth, Thomas A
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: Power-off tests were made in the 6- by 6-foot test section of the Langley stability tunnel to determine the variation of the static lateral stability characteristics with vertical-tail area, fuselage length, and wing dihedral. Two NACA 23012 rectangular wings with rounded tips and dihedral angels of 0° and 5° were tested alone and in combination with three circular fuselages of different lengths. The wing-fuselage combinations were tested as low-wing monoplanes with and without a horizontal tail and with variations in vertical-tail area. The results are presented as curves showing the variation of the static-lateral-stability slopes with angle of attack, and the rolling-moment, yawing-moment, and lateral-force coefficients with angle of yaw.
Bibliography:
Includes bibliographic references (p. 13).
Statement of Responsibility:
by Thomas A. Hollingworth.
General Note:
"Report no. L-8."
General Note:
"Originally issued April 1945 as Advance Restricted Report L5C13."
General Note:
"Report date April 1945."
General Note:
"NACA WARTIME REPORTS are reprints of papers originally issued to provide rapid distribution of advance research results to an authorized group requiring them for the war effort. They were previously held under a security status but are now unclassified. Some of these reports were not technically edited. All have been reproduced without change in order to expedite general distribution."

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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 - 003620415
oclc - 71360973
sobekcm - AA00006259_00001
System ID:
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Full Text






NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS





WARTIME REPORT
ORIGINALLY ISSUED
April 1945 as
Advance Restricted Report L5C13

INVESTIGATION OF EFFECT OF SIDESLIP ON
LATERAL STABILITY CHARACTERISTICS
II RECTANGULAR MINING O CIRCULAR FUBELAGE
WITH VARIATIONS IN VRTICAL-TAIL AREA AND
FUSELAGE LENGTH WIT AND WITHOUT
HRIZORHTAL TAIL SURFACE
By Thomas A. Hollingworth

Langley Memorial Aeronautical Laboratory
Langley Field, Va.






SWASHINGTON
/3 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.

.1-S DOCUMENTS DEPARTMENT
... .i ..



































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in 2011 with funding from
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NACA ARR No. LC13 RESTRICTED

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS


ADVANCE RESTRICTED REPORT


INVESTIGATION OF EFFECT OF SIDESLIP ON

LATERAL STABILITY CHARACTERISTICS

TI RECTANGULAR MTIDWING ON.! CIRCULAR FUSELAGE

WITH VARIATIONS IN VERTICAL-TAIL AREA AND

FUSELAGE LENGTH WITH AND WITHOUT

HORIZONTAL TAIL SURFACE

By Thomas A. Hollingworth


SUMMARY


Power-off tests were made in the 6- by 6-foot test
section of the Langley stability tunnel to determine the
variation of the static lateral stability characteristics
with vertical-tail area, fuselage length, and wing dihe-
dral. Two NACA 25012 rectangular wings with rounded tips
and dihedral angles of 00 and 50 were tested alone and
in combination with three circular fuselages of different
lengths. The wing-fuselage combinations were tested as
midwing monoplanes with and without a horizontal tail and
with variations in vertical-tail area. The results are
presented as curves showing the variation of the static-
lateral-stability slopes with angle of attack, and the
rolling-moment, yawing-moment, and lateral-force coeffi-
cients with angle of yaw.

The results indicated that the wing-fuselage inter-
ference on the slope of the curve of yawing-moment coef-
ficient against angle of yaw Cn, and on the slope of
the curve of lateral-force coefficient against angle of
yaw Cy was small and remained practically constant
over the installed angle-of-attack range. In the high-
speed flight range, the wing-fuselage interference on
the vertical tail was small and, in the normal flight
range, was not appreciably changed by an increase in
fuselage length or vertical-teil area for the sizes
investigated.
RESTRICTED









2 NACA ARR No. L5C13


'"ith no vertical tail, increased fuselage length
caused a negligible change in Cng for the fuselage
lengths tested. For the complete model, C increased
approximately linearly with fuselage length. The
slopes CnG, and 0 increased op;roxirately linearly
with vertical-tail area. For the system of axes used,
the slope of the curve of rollin,.--monment coefficient
against *angle of yaw C I, increased with vertical-tail
area at negative and small positive angles of attack but
decreased at large positive angles of' attack. The results
also indicated that increased dihedral angle slightly
decreased the rate of change of Cn with vertical-tail
area but had a negligible effect on the rate of change
of Cn, with fuselage length. An appreciable increase
in On. was caused by the end-plate effect of the
horizontal tail on the vertical tail.


IN TRODU CTI ON


The trend toward greater speed and higher wing loadings
and the increased consciousness of the importance of
satisfactory flying qualities have resulted in additional
attention being given to handlirg characteristics in air-
plane design. M!.'athematical eq4ations and convenient
charts for predicting the lateral stability characteristics
are given in reference 1. In order to use this material,
however, it is necessary to knov the stability derivatives,
which vary with each airplane configurat on. A series of
investi-ations has therefore been undertaken in the
Lanpley stability tunnel to determine the variation of
both the static-stability ani rotary-stability slopes with
various airplane parameters.

The present investigation is a continuation of the
investigation described in reference 2 cx.ept that, for
the present tests, the fuselage was equipped with a
rectangular wing in the midposition. The purpose of the
investiLation, vhich was conducted in the 6- by 6-foot
test section of the LIngley stEbility tunnel, was to
determine experimentally the effect, with the propeller
off, of vertical-tail area, fuselage length, wing dihedral,
interference, and the presence of the horizontal tail on









NACA ARR No. L5C13


the static lateral stability characteristics. P geomet-
rically similar model has beer, tested in the Lan3gley 7- by
10-foot tunnel (reference 3) and the data may be used to
correlate the results in the two tunnels.

Tests were made of a model that had dimensions pro-
portional to those of the average airplane. The ratios
of fuselage length to wing span and of vertical-tail
area to wing area investigated were taken to bracket the
range commonly used on present-day airplanes.


APPARATUS AND MODEL


The tests were made in the 6- by 6-foot closed-
throat test section (adjusted for straight flow) of the
Langley stability tunnel.

A three-view drawing of the model tested, which was
constructed of laminated mahogany, is given in figure 1.
Figure 2 shows the model mounted on the three support
struts for tests in the tunnel.

The two rectangular wings used for the tests have
dihedral angles of 0o and 5 and, in elevation, the
maximum upner-surfsae section ordinates are in one plane.
Each has an aspect ratio of 6.4 and an area of 361 square
inches, which includes the portion inside the fuselage.
The NACA 25012 profile is maintained along the entire
span.

The fuselage is of circular cross section and was
constructed as described in reference 2. Its dimensions
are presented in table I. With the shortest tail cone
attachl-d, the fuselage is geometrically similar to that
of reference 5-.

Five interchangeable vertical tails and the horizontal
tail were made to the NACA 0009 section (fig. 1). Their
dimensions are presented in table II.


TESTS


The wings with dihedral angles of 0 and 5 were
tested alone at angles of yaw of -50 and 5 over an









NACA ARR No. L5C13


angle-of-attack range from -100 to 200. The model combina-
tions tested are shown in table III. Model combinations
were tested at Pnr-les of yaw of -50, 00, and 50 over an
angle-of-attack range from -100 to 200 and at angles of
attack of -0.20 and 10.30 over an anrle-of-yaw range
from -300 to 120.

Tests in which the angle of attack was varied were
run at a dynamic pressure of 65 pounds per square foot,
which corresponds to a Reynolds number of approximately
888,000 based on an 8-inch wing chord. Tests in which
the angle of yaw was varied were run at a dynamic pressure
of 65 pounds per square foot at an arile of attack of -0.20
and at LO pounds per square foot, which corresponds to
a Reynolds number of about 546,000, at an angle of attack
of 10.30 to minimize the possibility of compressibility
effects at large angles of attack.

The rolling-moment data are not presented for a few
tests, because the tare readings were inconsistent.


PFESE;TTATION OF DATA

The results of the tests are presented in standard
NACA coefficients of forces and moments. Rolling-moment
and yawing-moment coefficients are given about the center-
of-gravity location shown in figure 1. The data are
referred to the stability axes, which are a system of
axes having their origin at the center of gravity Lnr.d in
which the Z-axis is in the plane of symmetry and perpen-
dicular to the relative wind, the X-axis is in the plane
of symmetry and perpendicular to the Z-axis, and the
Y-axis is perpendicular to the plane of symmetry.

The coefficients and symbols used are defined as
follows:

CL lift coefficient (L/qSw)

CD drag coefficient (D/qSw)
CY lateral-force coefficient (Y/qSw)

CVY slope of curve of lateral-force coefficient
against angle of yaw (6Cy/, )

CL rolling-moment coefficient (L'/qbSw)









NACA ARR No. L5C13


CZ slope of curve of rolling-moment coefficient
against angle of yaw 60C/6l)

C yawing-moment coefficient (I/qbSw)

Cn% slope of curve of yawing-moment coefficient
against angle of yaw c(6n/6 )

A1 increment of Cnw or Cy| caused by wing-fuselage
interference

A2 increment of C or Cy caused by wing-fuselage
interference on vertical tail

--S tail-volume coefficient
bSw
D force along X-axis; positive when directed downstream

Y force along Y-axis; positive when directed to the
ri ght

L force along Z-axis; positive when directed upward

N yawing moment about Z-axis; positive when tends to
retard right wing

L' rolling moment about X-axis; positive when tends
to depress right wing

q dynamic pressure I-pV2)

V free-stream velocity

P mass density of air

Sw wing area (2.507 sq ft)

b wing span (4 ft)

F dihedral angle, degrees

I tail length; measured from center of gravity, which
is assumed to be 10.40 inches behind nose of
model on center line of fuselage, to hinge line
of tail surface








NACA ArR No. L5C15


Sv vertical-tail area

a angle of attack, degrees

'V angle of yaw, degrees

The static-lateral-stability slopes Cn, C'',
and Cy. were obtained fr- 9 data measured at = +'
since the yaw tests showed that the coefficients hod an
a 2oximately linear variation in the range of angle of yaw
from5o to -5 In order to indicate the validity cf this
procedure, the slopes obtained ifrom yaw tests at = 0
are plotted with tailed symbols in the figures.

The accuracy of Cn, CG, and Cy was determined
experimentally to be about 0.0005, O.C'000, and
0.001, respectively, at a dynamic pressure of 65 pounds
oer square foot. The average experimental accuracy of
Cn C07, and Cy. is about 0.00010, 0.00016,
and 0.0002, respectively. The accuracies of the arple-
of-attack and an4le-of-yaw measurements are about 0.1-
and 0.050, respectively.

TAh.n of attacl- and dra- coefficient were corrected
for tunnel-wall effect by the following formulas:

Sw
Aa = 57.358w -OC =0.609CL (deg)

C 2 L

ACD = 6 C-L2 = 0.0106C2

where

6w jet-boundary correction factor at .:.ing (0.1525)

C cross-sectional area of tunnel (56 sq ft)

Both corrections are arditive. I'o jet-boundary correc-
tions were applied to C n, n, and Cy. The correction
to Cy is within the ey.oerimental error, whereas the
corrections to On and CO would be subtractive and
equal to about 1 percent.









NACA ARR No. L5315


The CL and CD data were corrected for the support-
strut effect; no corrections were applied to Cy, C ,
or Cn since previous results indicated the magnitude
of these corrections to be small for this model and support
system.

The values of L1 and A2 for C for the model
without wvina fillets were obtained by the following
formulas:

Cn =Cnwing-fuselage combination n wing +Cfuselage)



A2 n= Ccomplete model n*wing



+ Cn fuselage with hor. and vert. tails on C



The values of A1 and A2 for Cy may be obtained in
the same manner. The method used to obtain A1 and A2 is
the same as that of reference L. The following formula
(by which the value of Cn for the complete model is
obtained) is an example of the application of the incre-
ments A1 and A2:

S = C + C
Cn wing fuselage with hor. and vert. tails on


+ aICn* + A2Cn,


The interference between the fuselage and vertical
tail and the interference between the fuselage and
horizontal tail were not determined.

Lift-coefficient and drag-coefficient data for
representative model configurations are shown in figure 3.
The lateral-stability slopes Cn and Cy for the wing









:JACA ARR No. L5C13


are presented in figure 4.. The data presented in the
figures are summarized in table IV.


DISCUSSION


The static-lateral-stability slopes C and CY
remained practically constant over the installed angle-
of-attack range (figs. 9 to 15, 15, and 16). vTith the
system of axes used, the center of pressure of the vertical
tail varied with respect to the X-axis. At nei'ative and
small positive angles of attack, the center of pressure
was above the X-axis and, therefore, the side force on
the vertical tail caused a positive increment of C01,.
The opposite was true at large positive angles of attack,
since the center of pressure of the vertical tail was
below the X-axis.

The jar-. in the curves of lateral-force, rolling-
moment, and yawing-noment coefficients noted in figures 8,
15, 16, and 18 can probably be attributed to vertical-
tail stalling.


Interference Effects

The increments caused by v7in -faselzge interference A
and by ,'ring-fuselage interference on the vertical tail A
were computed by the equations previously given. The
fusela-e data (with and without tail surfaces) used in
these computations were taken from reference 2. The
other data were obtained from the present investigation.

The m:.-irnitudes of A10 Cn and Aicy0 are small and
remained .ratically constant over the installed angle-
of-attack r-i.ge (fig. 5). The chaige in the rragnitude
of these quantities with fusela-e lev gth was within the
experimental accuracy for the fuscl.0-e lengths tested.

Both A2Cn, and A2C0, varied a.'-precisbly with
angle of attack but their magnitudes were small in the
high-soeed flight rr.nge. (3ee f'i. 6.) -eplacin6
vertical tail 2 by vertical tail (a !S-percent









NACA ARR No. LC501


increase in area) orly slightly changed the magnitude of
these quantities in the normal flight range. As indicated
by previous experimental data (reference 4), the varia-
tion of 62(1n, and A20v with the fuselage lengths
tested was generally small in the normal flight range.


Effect of Horizontal Tall

Theory indicates that the presence of the horizontal
tail increases the effective aspect ratio of the vertical
tail and thus increases Cn and Cy (reference 5).

A pronounced increase in these quantities was obtained
in the present investigation by the addition of the
horizontal tail. This increase diminished somewhat with
a positive increase in angle of attack. (See figs. 7
and 8.) A correlation of the results of previous airfoil
tests in the Langley stability tunnel indicates a value
of 0.105 for the section lift-curve slope of an
NACA 0009 airfoil. by substituting this value for the
theoretical section lift-curve slope of 0.109 in equa-
tion (4) of reference 6 and by the use of figure 5 in
reference 5, an Incremental increase in Cy* of 0.0010 was
computed for the end-plate effect of the horizontal tail
on vertical tail 4. An average experimental increment
of 0.0010 was obtained for the model with a dihedral
angle of 0 and 0.0015 for the model with a dihedral angle
of 50 The end-late effect of the horizontal tail
on C amounted to about 10 of effective dihedral.

With the vertical tail off, the magnitude of the
static-lateral-stability slopes was not appreciably
affected by the addition of the horizontal tail. (See
figs. 9 and 11.)


Effect of Changes in Fuselage Length

Within the scope of the present investigation, a
negligible increase in Cn was obtained by increasing
the fuselage length for the model with no vertical tail.
(See figs. 9 to 11.) For the complete model equipped
with vertical tail 4., the increase in Cn with fuselage









NACA ARR No. L5C13


length was approximately linear An,' fairly constant over
the installed anle-of-attack ranl&e. (Fee figs. 12 to 14!.)
:he variation in C, and C, was small both with and
without a vertical tail for the fuselage lengths tested.


Effect of Chan-es in Vertical-Tail Area

The increases in Cn and Cy* with vertical-tail
area were approximately linear and the nmegnitudes were
nearly constant over the installed angle-of-attack range.
(gee figs. lh to 16.) As would be expected, at negative
and small positive angles of attack, Cz* increased with
vertical-tail area whereas, at large positive angles of
attack, C1 decreased with increased vertical-tail
area.


Effect of Changes with Constant Tail Volume

In figures 17 a;is 18 the result of changing the
fuselieI length and vertical-tail area in such a manner
as to hold the tail volume constant is shown. The configu-
rations tested in which the tail volume remained constant
are shown in table V. Data from figures 17 and 18 are
cross-plotted in figure 14. All the vertical tails tested
had an aspect ratio of 2.15.

The slope C,,, should remain approximately the
same vith constinJ"'tail volume. The small experimental
variation is possibly caused by interference or might be
explained by the arbitrary manner in which the tail-
volume coefficient was defined.

The values of CT and Cy~ are dependent mainly
on vertical-tail area and are practically Independent of
tail length (fig. 14). For the rrange of variations giving
constant tail volume, the change in C1. was not more
than about O.O00c2, which is equivalent to about 1 of
effective dihedral.








NACA ARR No. L5C15


Effect of Changes in Dihedral

The slope rC generally was slightly greater
for 5 = 50 than for P = 0. (See figs. 9 to 13, 15,
and 16.) W1th the vertical tail off, the change
in i with dihedral angle was insignificant (figs. 9
to 11) but, v;ith the vertical tail on, Cnp was slightly
larger for P = 0 than for F = o (figs. 12 to 16).
Figure 1L shows thpt increased dihedral angle slightly
decreased the rate of change of C with vertical-tail
area out had a negligible effect on the rate of change
of 'n with fuselage length.

The change with dihedral angle of wing-fuselage
interference and wing-fuselage interference on the vertical
tail was small.


Comparison of Data from Langley 7- by 10-Foot

and Langley Stability Tunnels

The model tested in the Langley stability tunnel
is 0.8 as large and geometrically similar to the one
tested in the Langley 7- by 10-foot tunnel for the
investigation of reference 3. The test Reynolds number,
based on the wing chord, was about 619,000 for the
Langley 7- by 10-foot tunnel compared with about
888,000 for the Langley stability tunnel. The effective
Reynolds number, however, was about the sarre since the
turbulence factor for the LanLley 7- by 10-foot tunnel
is 1.6 compared with less than 1.1 for the Langley stability
tunnel. Data taken from reference 5 were converted to
the stability axes and the angle of attack was corrected
for tunnel-wall effect in order to make the data comparable
with data from the Langley stability tunnel. Figure 19
shows that satisfactory agreement was obtained for all
three static-lateral-stability slopes. In both tunnels
the model, when yawed, tended to roll violently at the
stall.

CONCLUSIONS

The results of tests of a model consisting of a
rectangular midwing on a circular fuselage with variations








NACA AHR Ho. LSC1d


in vertical-tail area and fuselage length with and with-
out a horizontal tail indic: ted., for the ran-e of con-
figurations tested, the following conclusions:

1. The wing-fuselaLe interference on the slope of
the curve of yawlng-moment coefficient against -tn le of
yaw Cn* and the slope of the curve of lateral-force
coefficient against angle of yaw Cy was call and
remained practically constant over the unstrlled .tngle-
of-attack range. In the high-speed flight range, the
wing-fuselage interference on the vertical tail wa5
small and, in the normal flight rsnie, was not appreciably
changed by fuselage length or by an increase of about
.8 percent in vertical-tail area.

2. The end-plate effect of the horizontal t-il on
the vertical tail appreciably increased Cn., and Cy,,.
Good agreement was obtained between experimn.-nt.al and
computed values of Cy .

5. Increasing the fusel'.-we length with no vert-cjl
tail resulted in a negligible change in Cn for tl-.e
model, both with and without a horizontal tail. ,or the
complete r-odel, the increase in Cn, was approximately
linear with fuselage length. The changes in th.e. slope
of the curve of rolling-moment coefficient against angle
of yaw Ci, and in C with fuselage length vere-
small.

4. The increases in Cn and Cy with vertical-
tail area were approximately linear. For the system
of axes used, an increase in vertical-tail area increased
0C at negative and small positive angles of attack Lut the
opposite was true at large positive angles of attack.

5. Tncrpased dihedral an: .e slih-tly de,'reaser' the
rate of change of C with vertical-tail 7'rea but had
a nelt-i-ble effect on the rate of ch... of n, ilth
fuselage length.

LDn;:le:; memorial Aeronautical Laboratory
nationall Advisor'; Committee for Aeronautics
L-imgley field, Va.








TV.CA ARR I-o. LC13


REF5cE >'.CES


1. Zimmerman, Charles H.: An Analy is of Lateral
Stability in Power-Cff EliCht v.ith Churts for Use
in Des_;n. IACA Ren. N:. 53?, 1937.

2. Fehlner, Leo F., and "'scLachlan, Fobert: T nvesti-
gation of Effect of SIi.esli' on Lateral St. ilit
Characteristi cs. f "ircular F.JselEr-: wit;i
Variations in Verti.c'l-Tril Area and Tail Length
with and 'Aithout "'oriz-ntal Tail Surface. iACA
ARm No. LEE25, lt1.

5. Bamber, M.'. J. and House, '. 0.: Tind-Tunnel Investi-
gation of Effect of Yaw on Lateral-Stab4 lity
Characteristics. TT Rectangular Nr.A.C.A. 25012 Wing
with a Circular Tuselage r.n a Fin. IJACA mIT "3o. 730,
1o3?.

4. Pecant, Isidore O., and 7.!allsce, Arthur R.: lind-
Tunnel Tnvestigation of Effect of Yav on Lateral-
Stability Characteristics. II Symmetrically
Tapered ?'inr at Various Positions on Circular
-uselage with and without a ',ertical Tail. :'ACA
TN Fo. 825, 19L1.

5. Katzoff, S. and r'utterneri, 7Tilliaim: The End-Plate
Effect of a Forizontal-'lail S.1rface cr a Vertical-
Tail Surfaze. IACT A T I.;'. 7`7, 19L1.

6. Jones, Robert T.: Theoreticc:. Corre-cton for the
Lift of Tll1 tic "irn:s. Jour. Aero. S, .1 vol. 9,
no. 1, ov. 19:l, po. -O10.









NACA ARR 1To. L5C13


TABLE I

FUSELAGE DIKENSIONS


TABLE II


TAIL-SUPFACE DIT7E'SIONS


Area measured from root shord at center line of fuselage.

NATIONAL ADVISORY
CO'M-TTEE FOR AERONAUTICS


F1'selege Tsil-cone Tail length, I Tall length I
Fuselage length length (in.) '
(in.) (in.) -E soan

Short 32.25 9.`5 20.07 0. 18

.edia u 57.05 1. 65 2h-7 *512

Long '1.65 190.L5 29.67 .618









5 A.1.:, ARR .'o. L C.L ,





TABLE III

MODEL COMBINATIONS TESTED

[r = 00 and 50]

Horizontal Vertical Fuselage Variable
tail tail Fuselage Variable

Off

1

2 Short,
medium, a
5 and long

4

5
On
2
Long
4

-- 1'edi um


Short
Off

Off Off Long a and 4

4 Short


I ATITI AL ADVISORY
CO01SJTTEE FOR AEPOINAUTICS







NACA ARR No. L5C13


TABLE TV

PR'7EmV'TATIOU OF RESULTS


Figure Descriotion of figure Data
n:presented


5


h



5


6



7


8


9


10


11



12



13


Effect of changing fuselage length
(horizontal tail and vertical
tai 1 o n) .- _1


C0 and
CD as

ln and
C as
.Y'!


Aln$ and
61Y as

62Cnw and
62C as


f(a)


f(a)


f(a)


Lift and drag curves for repre-
sentative model configuration

Slope of yawing-moment and
lateral-force coefficients for
NACA 25012 rectangular wing

Effect of wing-fuselage inter-
ference

Effect of wing-fuselage inter-
ference on vertical tril

End-plate effect of horizontal
tail


End-olate effect of horizontal
tail

Effect of changing fuselage
length (no tail surfaces)


Effect of changing fuselage length
(no tail surfaces)

Effect of changing fuselage length
(horizontal tail on; vertical
tail off)

Effect of changing fuselage length
(horizontal tail and vertical
tail 4 on)


Cn, C ,
Gy as


and
f l )


'n*, C and
,y as f(a)


rn r ,
Qy as


Cn,
CY41


as


and
f(W)

and
f(a)


Cngr CjZ, and
Cy as f(a)

Cn, -IL and
Cy as f(,)


NATIONAL ADVISORY
COMTYTTEE FOF AERONAUTICS


Cnw, C0e, and
C. as f(a)


I








I-ACA ARR No. L5C13


TA3'S IV Concluded

7PRESENTATI OF,? RESULTS Concluded


Figure Description of figure Data
presented

1i. Effect of cha',. -ing fusell,.-e length Cn1, Cj ,, and
Cy as f

15 Effect of changing vertical-tail Cn,, 0C, and
area v
Cy. as f(a)

16 Effect of chn t -in vertical-tail Cn, CZ, and
area Cy as f(*)

17 Effect of changes with tail volume C CC1, and
constant v
Cy as f(a)

18 Effect of chL.ngs with tail volume Cn, CL, and
constant C as f()

19 Crmparison of data fr.. T in,..;.ley Cn., CL and

10-foot tu'nels C, as f(a)


NATIONALL ADVISO-.Y
CO"IiTTEE FOR ,...'.-ONAUTICS









NACA ARR No. L5C15


,: I ,D




Cp
C






Fa

C
I








r4
-4'














IC ,







E->i ,O
E2K

0






'-4
El-.

co
(a<
I.,


.- C a -,
rI ['- o
0 0 0


LO CD CO
,-1 1-1 1-4
zt LC s
0


o









Sr-r c 1


0
o
E-
cc.






c P


-0


o o o
0 0 0
0







NACA ARR No. L5C13


FUSEL AGE


TAIL 3
r 128 R





TAIL 4


Fig. 1

4 7
Il .


TAIL 5


TAIL 2


A AND B ARC QUADRANTS
OF SIMILAR ELLIPSES


-------- : ':- 5 -"


NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS


I'




4185~zi


Figure I Rec angular NAACA 230/2 w,r g in comnbnoaon ,v th
circular fuse/age, ver/c al ando horizontal tods, and
talo cones All dimensions given in inrces.


32.25-4 37.5--4









NACA ARR No. L5C13 Fig. 2






-o



Cd

c-

D -







c ,D










0)
lid id






CO


..-4 (








r" 0
cJ
'l .-$ w








cd 0
GiE





OD .L-






-4

bio









NACA ARR No. L5C13
















it? -- ---













.4 -

U-






O -- --


-12 -8 -4 0 4 /12. 16 20

Angle of oftock, oC deg

Fgu/we 3.- Va'oro/7 of ///t om/ d-rafy ceff/'ielts with anf/e of
af/fc/r for reptesenfatIve ae/ ~ eol ur/oons f, I"
6ff i// ft/ *


Fig. 3





Fig. 4 NACA ARR No. L5C13





2c
X

0i0
_-K- ----- _

-_____ -5 __ ^










N N _', ', _I _
---it----1-- ^





th
A
I I tI

I'- __ ----- __


(SN
-- -- ---_ _-^ -- INN






NACA ARR No. L5C13 Fig. 5









q-j


OO..
,-- .

-- C o .













19C
op
K .,




--- I --I -- -- -- QJ -




.^'1
___ / ___ ___ ___ ___ __ ;| ___ ___ ___ _i N
aS K


C
- N






NACA ARR No. L5C13


ox+o
_k & -- -- -- -- K > -- -*-(


















o o
S II
------ t- -



























. |"
0 ?
___J 0 ^


N ---- < ^





_-"-- 3 0

---- I---- >




I '^



(4 K


Fig. 6a







NACA ARR No. L5C13 Fig. 6b






.Q















-4 C




-I- -J x"
lba










t J__ __ [_






0 0 1% -
--,--,- ^ I

(3V^
--- -- --- ( -- -- --- --- -- a






Fig. 7a


NACA ARR No. L5C13





1--
K
x








o, "



- ~)S
\ -\
r

~13

f- Q1^

- 1'jri

/ l i
x i




4 1r4
^ N Li


AC







NACA ARR No. L5C13


Fig. 7b


a
k














I'

K





NACA ARR No. L5C13


04 Horizonta/ Vertical
f aol/ ta"d
x Off Off
.O o Off 4
+ On 4

















An~/e of yaw 1, deg


&)ure 86 Ef4ect of horizontal /oa/ urface on
variotion of yow/y-mornment, rollh,-moment,
and /afero/-force coeffi'rn.r wi/h on/le oat
yow. Short t%,Je/oge wi/h vertfcoa/ *a/ 4.
_^:^:3^?F_7


-^ ^---70 i


Fig. 8a







NACA ARR No. L5C13


.04















Ni,-^3 ~--------
S- x


























Angle of yaw
0r


















--30e -2 -/0



66) 1 ; cr 10.3
F & Contaed.


Vert/c e/
toa/


'1% s
tz^

0It

z'h


'- deeg

; a 40 /b/.rg ft .


Fig. 8b






NACA ARR No. L5C13


H/-or/zon tO/ Verficol



ox Off 4










. .










-0 -04




RATIONA ADVIS( RY
COM ITTEE F)R AERO IAUTICS

-00 -20 -/0 0 /0

Angl/e of yaw ,*r deg

[ 5 a I -0 -.2 6 q 651/,/q ft.
F0,ure 6 -. Co ---nued.
-- -- -- -- -- -- .0


Fig. 8c






NACA ARR No. L5C13











I 0














./0




>> ~ KT


-.20


-30 -20 -/0 0 /0


Angle of yaw


, V de0g


(d) r
fflgure 8


, S 5 a /0.3
. Conc/ueo' .


9 40 /b/sq ft .


Fig. 8d






NACA ARR No. L5C13


co
"s..

















co

I


I .
1%,

%
t\







AL








z
%
S10

b .'<"
N i


? s
,


(
Ak


Fig. 9






NACA ARR No. L5C13


k02
A.


















4/0


-30


-20 -/0 0
An/gle of yaw p jy d{p'


Po) 0 0 a., -0-.2 0 654 /j/sq ft.,

Figufre /0. Effecf of eha'nfi futre/aye /e/ yth on
vor/at/on of yovw/7n-momen7;, ro//ly-moment,
an/7 /atro/-force coeff/ents with ong/le of
yow HorzonfoA/ and verhe/a/ /a//p off.


Fuse/oge


o Short
+ Long




NATION L ADVI RY
CON MITTEE OR AERC IAUTICS

_ _+ -4e-r


Fig. lOa


is







NACA ARR No. L5C13


.02


I /O
N-0
g%
-1


", OZ


.+----4


Fuse /age
.02

o Short
+ Long -






-.02+







NATION ADVISORY
CO IMITTEE -OR AERONAUTICS


-30


-2z0


-/0


Angle of yaw ,


k -, deg


(b) /" 0 ac 10.3 ; 9 40 /b/1f ft.
Figure /O Continued .


Fig. lOb






NACA ARR No. L5C13


.02


O
1%












thj






/0



Jo
N^02
No -/


-30


I3
q)

L N
OO


-2O -/0 0 /O.

Ang/l of yow *t deg


fc) /P, 5 ; i .
Figure /0 Continued .


; 9 665 /I/srq ft.


Fig. lOc






Fig. lOd



02




















do
>-













1.


tq o
K N


-o
k 'S
Q., S$9


NACA ARR No. L5C13















- .02
O




q1
z3 $
P


-30 -20 -/0 0 /0


Angle of


6o)
F/,ure /O.


Ye7w,3,dg


, 40 /bhq ft.


/ d; a ;
- Concluded .





NACA ARR No. L5C13


b
O
'II
O


0
4!







OO
z44
S 0q
y !!<



U oj>
'I"-j

*% -.i
Iz -(


E~4


Fig. lla






NACA ARR No. L5C13


0


I'





-Q




I3


to








I
!5


^


L


Fig. llb







NACA ARR No. L5C13 Fig. 12














SI






















0 + 0 0 -
q ^ & ,
1 i -S~z C \ 0k
_^-<^ _, .--sL- y .
^ *S l ^
-^^4.-^zW^ -- ^i


hL
13






NACA ARR No. L5C13


.04


2o02




-.OZ












'' N NAION& NISO




-$O -20 -/0 0 /o
A,-g/e of Y w W o'g


Fi-ure 1/.-Eff'c of chnaginy fue/ay6 lengthh on
variot/on of yawrng-moment ro///mn-momenrt,
4nd /uWra/-force coeff7cientr with ang/e of yaw.
//orrzon01o/ to/0 on'7d ve-rtcal tafoll 4 on.


Fig. 13a







NACA ARR No. L5C13






.04




S".02 -


0





O -






./O -
./0




.O --




t0



-. 2 O


-JO


Fig. 13b


-20 -/0 0 /0
Angle of yaw ), oCey


fb6) r, 0 ; ar /0.J ; q 40 /b/1 ft .
/ggure /J Continued .






NACA ARR No. L5C13


Ok + + Medium
02 ---- --x Long







..02
-^0 ----- ------ ---


















0-- -.04



.. ........... A
1-0 -20 -/0 0 /0

Angl/e of yow, Y, "o'e

c() /3 5 -0.2 // f
E 3g0re / Con/inued .


Fig. 13e







NACA ARR No. L5C13


.04



.02







-.02


./0


z "
, Qj
?(1
r


S1j
fS


-30 -z. -/O 0 /0
Ang/e of yaw 2 o eg

(o') F/ 5. /0 /O.J; 40 /b/q f/ .
F/,gure /J Conc/uoe'd.


Fig. 13d






NACA ARR No. L5C13


Dihedrol
(deg)
----5


---0


(,"?edra/
(deg)


o Sort 0
+ Med/,am 0
x Lony 0
a Jhor/ 5
o A4ed/umn 5
A Lon i


0 .04 .08 .12 ./l


fo) 0 *.
F/ire / 14.-Effect of chang/n1 fuse/aqe length
on variation of /aleral-sfabi'ly slopes Cl,,,,
Cz anod Cy/ with vertical/- a// area.
Horizontal/ //il on ; q 655 / tb/ y /).


el7p


Fig. 14a







NACA ARR No. L5C13


Dihedral
(deg)
- -- -- -- 3


0 -


COX-

".4








76-


Fue/lage


Short
MedOam
Log
Short
MAedurn


0 .04 .08



() cr /0.


Figure /4 Concluoeod.


i/e dra/
(dey)

0
0
0
5
5
S


./2


Fig. 14b






NACA ARR No. L5C13


t j



L-Z


Fig. 15a






NACA ARR No. L5C13


Fig. 15b





%














( O
0.

L*
Co







NACA ARR No. L5C13


.04 -1- III rol/


o__ o
+ V4
02 __ __






















-30 -20 -/0 0 /-
-.O0 -.02

















Angle of yaw deg

to) 0/, ",-0.2* ,6/1b/lrqft.
Figure /6 Effect of changing vertical a;l
area on0 varia'/ion of yawing momet/ ,
rollingg moment and a/eral force
coe efficient wi/h a7nge of yaw L on7
fue/alage with horIzontla/ ol/


Fig. 16a







NACA ARR No. L5C13


.04 *-I ver









0




-.02












-o






h "I NAT FINAL AE VISORY
-- -- fMJ-LEL














14-1 I4MMITT F FOR A :RoAITi CS


-30 -20 -/0 0 /0
An,/e of yow, ya 7lk

<) /P. o, 0 /0. 4,/ 40 //. .
F/.reo /6.- Con--------


Fig. 16b






NACA ARR No. L5C13


.04







-.02











./0









-20


-30


-20
Angle


-/0
yaw


0
Y ,


F f)re /6 *; d 6 //
/I-gure /6 Confmateo' .


Fig. 16c


Sor

I .6
. _







NACA ARR No. L5C13


.04



I I
02















./0


s*2

L -. /
to
0
-.I0


-JO


-20 -/0
Angle of yaw


&d) / 5 ; a- /0.4 ;
Figure /6& Concluded .


Fig. 16d

7/



















1





.7


0 /0
- doey

q 40 /lb/lsq ft.






Fig. 17a NACA ARR No. L5C13




1/ o~










V~ -t
i \1 \












LL
L__z___ __-_ o I 6









l b
j_ \, 8










oo o o'
--C, L ',
^ ^_' ~ __511^ ^__\ '4b
0^ + 4 ^6 -7 3 >^
T1^^^ \^ 71 ^S'




_x. 1 q)_^ "* K ot o







NACA ARR No. L5C15 F'ig.


Ie

\, \I
*




N





-0 a 0


,\ o.
-
4' 3 o' S "o



-K--- V --t?-I








0+ x




/I o







-- --
--_ < --- -- _<: bSS

__ -





__ ___ __ __ __ */1_
N K
o o
------------C--------------


17b







NACA ARR No. L5C13'


>jII__


__1 ___~





- ___I ___


t 11 I 1 t~*~


C NffTIONAL ADVISO1
S_ COMM ITEE FO t AERON UTICS
,t


-20 -/0
Ang/e of yaw 7


fa(o) f 0 o -02 ; 65/bh/i ff' .
//iure / '. Eff/'ect of .eve'ro/ combo/,,ans
having co/n'/.a'n/ /,17o vo/l/'ne on varo'//ion of
yaw/.g, -mome n', ro/h/ig -,omnet and /a/ero,/-
force' coeff/'eA.i/'/, wi/h ay/7/e of yaw. To//
vo/ume 0. 0407 ; hor/zono'l/ /ail on .


Fuse/oge


o Short
+ Me diumn


1/6
7

]


,rtiao /


4
3








SL
Q'




-.02 ,


3i


01?
S0


.10


/0




-./0


0
3 deg


Sp


A0


Fig. 18a







NACA ARR No. LEC13


.04














.02
-.OZ


S.
~aC
1.~
'a

S..
~1



.20


-30


-20
Ar ,,9/


-/0 0
of yaw k# 4e9


Figure /8. Cont//ced


f6) /'', s; a ,/10.3; 0 ., 40 /I/h g ft.


Fig. 18b






Fig. 18c


.04


".02.
o

I N
.o


_0>.






.0





./O


to
(/O
i o


so
,^ ^-/
'a


NACA ARR No. L5C13


Fuse/og e V1e,r-Gico
toil

Shor9- 4
Med/urn 3
Long l



-N





__ .02




0 ,
.o


-2O -20 -/0 0 /0
Angle of yaw doeg
/C) /- 5 ; -0. 0 ; 2 65 /b/bq ft
,7.qure 18. Cont,'nueo .







NACA ARR No. L5C13


Fuse/ oge


.04



.02










./0
0


.02


%
-0
-.


O
. O
-.-/o


-o

-^2


Ve rti/o
to//

4
3
'a,







-2-


.02



- ,


-oV

A)


-20
Angle


-/0
of yaw


(d) r 5 ix /O.J r ; 40
Figure /6' Conc/ued .


/b/rq ft .


-JO


0
, P ,


d'eg


Fig. 18d






NACA ARR No. L5C13


s^1 K J



T~~~~ q' 1> ^L<

_4-_i __^ ^ ^-L


i^"-ro.0
N. f
^ N.
qj -i .




,-

q-s^^
1Z3 kt
'N Ot
*.- < o
t3 6 5
< 0.






It -
^ 'U s
to h^

.1*2 ^.
N. 5


Fig. 19








UNIVERSITY OF FLORIDA
lil 1 111 1111111111i IIII
3 1262 08104 972 7


iJNIVERSrPT OF FLORIDA
DOCUMENTS DEPARTMENT
120 MARSTON SCIENCE LIBRARY
F:O. BOX 117011
CA.NESVILLE,,FL 32611-7011 USA


A





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