Investigation of effect of sideslip on lateral stability characteristics

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Material Information

Title:
Investigation of effect of sideslip on lateral stability characteristics
Alternate Title:
NACA wartime reports
Physical Description:
20, 31 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 -- 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. 15).
Statement of Responsibility:
by Thomas A. Hollingworth.
General Note:
"Report no. L-17."
General Note:
"Originally issued April 1945 as Advance Restricted Report L5C13a."
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
Rights Management:
All applicable rights reserved by the source institution and holding location.
Resource Identifier:
aleph - 003617744
oclc - 71354417
sobekcm - AA00006260_00001
System ID:
AA00006260:00001

Full Text







NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS






WAllTIME REPORT
ORIGINALLY ISSUED
April 1945 as
Advance Restricted Report L5C13a

INVESTIGATION OF EFFECT OF SIDESL.P ON
LATERAL STABILITY CHARACTERISTICS
III REOTANGLmAR LOA WING ON CIRCULAR FUSEAGE
WITH VARIATIONS IN VERTICAL-TAIL AREA AND
FUSAGE LEGT WITH AND WITHOUT
HORIZONTAL TAIL SURFACE
By Thomas A. Hollingworth

Langley Memorial Aeronautical Laboratory
Langley Field, Va.






5V


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.


L-17


DOCUMENTS DEPARTMENT


--I-r L-/ 7

LI .-i


It


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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/investigationofe










NACA ARR No. L5Cla RESTRICTED

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS


ADVANCE RESTRICTED REPORT


INVESTIGATION ?'F EFFECT OF SIDESLIP ON

LATERAL STABILITY CHARACTERISTICS

III RECTANGULAR LOW WING ON CIRCULAR FUSELnaG

WITH VARIATIONS III VERTICAL-TAIL AREA Afl

FUSELAGE LENGTH WITH A. D '.WITHOUT

H.ORIZONTAL TAIL SURFACE

By Th.n-mas A. Hl)1lingrcrth


SU L i! AR Y


Power-off tests were rna.:.e in the 6- by 6-foot test
section of the Langley stability tunnel to determine the
variation of the static later-al stability characteristics
with vertical-tail area, fuselage length, and vin drbihe-
cral. Two NACA 5'3012 rectangular winL-s with riunoed tios
and dihedral angles of 0 and 5c were tested alone and
in combination with three circ.-labr fucelages of different
lengths, The wing-fuselage cormbinations 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, rawinJ-moment, and lateral-force
coefficients with angle of yaw.

The results indicated that the influence of wing-
fuselage interference on the slope of the curve of yawing-
moment coefficient against angle of yaw Cnw, and the
slope of the curve of lateral-force coefficient against
angle of yaw Cyv was usually stabilizing, appreciable,
and varied with angle of attack. The influence of the
wing-fuselage interference on the vertical tail was also
generally stabilizing and appreciable at negative and
small positive angles of attack but varied with angle of
attack.


RESTRICTED









,ACA ARR No. L5Cl5a


With no vertical tail, increased fuselage length
ordinarily caused a slight increase in Cn,, for the
fuselage lengths tested. At the larger negative angles
of attack, this effect was more pronounced. For tne
complete model, the increase in Cn was approximately
linear with fuselage length. The marniitude of this
increase appreciably diminished with a positive increase
in angle of attack. The slopes Cn.& and Cy,,, increased
approximately linearly with vertical-tail area. For the
system of axes used, the slope of the curve of rolling-
moment coefficient against angle of yaw Cj,, increased
with vertical-tail area at negative Z:-nd si.-,all positive
angles of attack but the opposite w--s true 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 cring.e of Cn,, with
fuselage length. Except at large positive angles of
attack, Cvy, was generally i.I'ater with the smaller
dihedral angle. A slight increase in Cni, -vas caused
by the end-plate effect of the horizontal tail on the
vertical tail.


INTRODUCTI ?U


The trend toward greater speed and nigher wing
loadingW and increased consciousness of the importance
of satisfactory flying qualities have resulted in addi-
tional attention being given to handling characteristics
in airplane design. Mathematical equations and conven-
ient charts for Dredicting the lateral stability charac-
teristics are given in reference 1. In order to. use
this material, however, it is necessary to know the
stability derivatives, which vary with each airPlane
configuration. A series of investigations has therefore
been undertaken in the Langley stability tunnel to
determine the variation of both the static-stability and
rotary-stability slopes with various airplane parameters.

The present investigation is a continuation of the
investigations described in references 2 and 3 exceot
that, for the present tests, the fuselage was equipped









NACA ARR No. L5Ci5a


with a rectangular wing in the low position. The purpose
of the investigation, which was conducted in the 6- by
6-foot test section of the Langley stability tunnel, was
to determine experimentally the effect, with the propeller
off, of vertical-tail area, fuselage length, wing dihe-
dral, interference, and the presence of the horizontal
tail on the static lateral stability characteristics. A
geometrically similar model has been tested in the
Langley 7- by 10-foot tunnel (reference 4) and the data
may be used to correlate the results in the two tunnels.

Tests were made of a model that had dimensions
proportional to those of the average airplane. The ratios
of fuselage length to wing span and of vertical-tail
area to wing area investigated ;v.ei-e 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 lam inated rrahogany,
is given in figure 1. Figure 2 shows the model r.-ounted
on the three support struts for tests in the tu.'-ncl.

The two rectangular wings used for the tesz9 have
dihedral angles of O and 5 and, in elevation, tis
maximTLT upp2r-su'r.t'ace section orcinates are in cre lane.
Each has .:in aspect ratio of 6.14 and an area Df 5(,61 square
incnes, which includes the portion inside the fuselage.
The IACA 25012 profile is maintained along the entire
span.

The fuselage is of circular cross section and was
constructed as described in reference 2. Tts dimensions
are presg;ntea in table I. With the shortest tail cone
attacherid, ihe fuselage is geometrically similar to that
of reference 4.

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









NACA ARR 11o. L5C13a


TESTS

The wings with dihedral angles of 00 and 5 were
tested alone at angles of yaw of -50 and 50 over an
angle-of-attack ran r from -100 to 200. The model combi-
nations tested are shown in table III. i.Iodel combinations
were tested at angles of ycw of -50, 00, and 50 over an
angle-of-attack range from -100 to 200 and at an ungle
of attack of 10.20 over an angle-of-yaw range from -500
to 120.

All tests were run at a dyn'mic pressure of 65 pounds
per square foot, which corresponds to a test Reynolds
number of approximately 888,000 based on an 8-inch wing
chord. The data may have been affected by compressibility
at large angles of attack.


FRESEi!TATION OF DATA

The results of the tests are ore.sented 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 a system of axes in which the Z-axis is in
the plane of symmetry and perpendicular to the relative
wind, the X-axis is in the plane of symn-ii.try and perpen-
dicular to the Z-axis, and the Y-axis is perpen.dicular
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/qS-,)

CyV slope of curve of lateral-force coefficient
against ingle of yaw (cCY/6 )

CZ rollin-moment coefficient (L'/qbS,)









NACA ARR


C L,


Cn
Cn



A1


A2


ISy
bSw

D


Y


L


N


L'


q

V

P

Sw
b

r


No. L5Cl5a 5


slope of curve of rolling-moment coefficient
against angle of yaw (C6/c4)

yawing-moment coefficient (rr/'qbs3w)

slope of curve of yawing-'moment coefficient
against angle of yaw (OCn/5 )

increment of Cnp or Cyp' caused by wing-
fuselage interference

increment of Cnz, .r Cy, caused by wing-
fuselage interference )n vertical tail

tail-volume coefficient

force along X-ax:s; positive when directed down-
stream

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

force along Z-axis; positive when directed
upward

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

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

dynamic pressure PV2)

free-stream velocity

mass density of air

wing area (2.507 sq ft)

wing span (4 ft)

dihedral angle, degrees









NACA ARR No. L5C13a


1 tail length; measured from center of gravity,
which is assumed to be 10.o0 inches behind
nose of model on center line of fuselage, to
hinge line of tail surfaces

Sy vertical-tail area
a ancle of attack, degrees

w angle of yaw, degrees

The static-lateral-stability slopes Cn,, Cl,,
and CVY| were obtained from data measured at & = +50
since the yaw tests showed that the coefficients had an
approximately linear variation in the range of angle of
yaw from 50 to -50. In order to indicate the validity
of this procedure, the slopes obtained :r!m yaw tests
at = 00 are plotted with tailed sv.rbols in the
figures.

The accuracy of Cn, C1, and Cy was determined
experimentally to be about 0.0005, 0.0008, and 0.001,
respectively. The average experimental accuracy of Cnt,
CL, and Cy* is about 0.00010, 0.00016, and 0.0002,
respectively. The accuracies of the angle-of-attack and
angle-of-yaw measurements are about 0.10 and 0.050,
respectively.

Angle of attack and drag coefficient were corrected
for tunnel-wall effect by the following formulas;

Sw
La = 57.356w- CL = 0.609CL (deg)


ACD = 6w -CL2 = 0.0106CL2
C

where

6w jet-boundary correction factor at wing (0.1525)


cross-sectional area of tunnel (36 sq ft)









NACA ARR No. L5C15a


Both corrections are additive. No jet-boundary correc-
tions were applied to CL. Cn, and Cy. The correction
to Cy is within the experimental error, whereas the
corrections to Cn and C w'.ould oe subtractive and equal
to about 1 percent.

The CL and CD data Aere corrected for the support-
strut effect; no corrections were ano:lied to Cv. CL,
or Cn since previous results indicated the magnitude
of these corrections to be snrail. for this model and
supoort system.

The values of L1 and A2 for Cng for the nodel
without wing fillets were obtained by the following
formulas:

AwC Cning-fuselage combination /C wing fuselage

% complete model \ ,.ing

+ Cn fuselage with hor. and vert. tails on ICn-)

The values of Ai and A2 for Cy| may be obtained in
the same manner. The method used in this investigation
to obtain A1 and A2 is the same as that of reference 5.
The following formula (by :which the value of Cno for
the complete model is obtained) is an example of the
application of the increments A1 and A2:

Cnlw = Cni + Cry
wing fuselage with hor. and vert. tails on

+ AlCn1 + t2Cry


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 5.
The lateral-stability slopes Cn1 and Cy, for the wing









FACA ARR No. L5Cl3a


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


DISCUSSION

For the complete model, the static-lateral-stability
slopes Cn, and Cy| usually decreased with a positive
increase in angle of attack (figs. 12 and 15). The
results of the present investigation indicate that this
decrease was caused by interference (figs. 5 and 6).
With the vertical tail off, the variation of these values
with angle of attack was irregular anrnarently also because
of interference (figs. 9 and 11). Such variations with a
of the lateral-stability slopes as were obtained in the
present investigation for the low-wing model both with
and without a vertical tail were not shown in the midwing
investigation (reference 3). The slopes Cy1 and Cn,
were practically always greater for the low-wing than
for the midwing configuration, apparently because of a
change in interference with wing location.

At negative and sometimes at small oositivc angles
of attack, CZ decreased as the angle of attack became
less negative (figs. 12 and 15). In the positive angle-
of-attack range, CL, generally increased with angle of
attack. The slope C, was increased because of the
side force on the vertical tail at negative and small
positive angles, of attack but the op'nosite was true at
large positive angles of attack. This effect may be
attributed to the system of axes used. For this system
of axes, the center of pressure of the vertical tail is
above the X-axis at negative and small positive angles
of attack; consequently the side force on the vertical
tail caused a positive increment of C1.. The opposite
was true at large positive angles of attack because the
center of pressure was below the X-axis. The slope C1
was appreciably greater for the midwing configuration
than for the low-wing configuration, frequently by as
much as 30 of effective dihedral (reference 5). This
change in slope is evidently caused by a change in the
nature of the flow around the wing near the wing-fuselage
juncture.









NACA ARR No. L5C15a


Interference Effects

The Increments caused by wing-fuselage inter-
ference Al and by wing-fuselage interference on the
vertical tail t2 were computed by the equations
previously given. The fuselage data (with and without
tail surfaces) used in these computations were taken from
reference 2. The other data were obLained from the
present investigation.

The quantities IlCn,,. and AiCy,, were generally
appreciable and had a stabilizing effect on the model
(fig. 5). The var"atLion of these values with angle of
attack was irregular,. but A3C Cv, generally tended to
decrease with a positive incrr.ase in angle of attack.
The irregularity of the curves m.ay :ob caliced by a burble
at the juncture of the wing and fuselage, additional
evidence of which may be seen in tne curves of lift and
drag coefficients in figure 35. An an.preciabl_ part of
the value of Cy| for the wing-fuselage combination can
be attributed to interference. The changes in 6lCn,i
and AiCy, with fuselage lengthwc.rewithin the experi-
mental accuracy for the fuselage lengths tested.

At negative and small positive angles of attack, the
quantities &2Cnw and A2Cy1, were gEnerally appreciable
and had a stabilizing effect on the iiooel (fig. 6). ldith
an additional positive increase in angle of attack, the
values changed in such a manner as to become dectabilizinp.
The effect of replacing vertical tail 2 by vertical tail b4
(a h8-percent increase in area) on these quantities was
generally small in the installed range. The variations
of A2Cnt and A2C'y with fuselage length nere somewhat
irregular. Because the model t,:ted in this investigation
had no wing fillet, caution should be used in applying
the results to design since the presence of a fillet may
appreciably change the lateral stability characteristics.
In view of this fact, an investigation of the lateral
stability characteristics of a model vwith wing fillets
might be desirable.









NACA ARR No. L5C15a


Effect of Horizontal Tail

Theory indicates that the presence of the horizontal
tail would increase the effective aspect ratio of the
vertical tail and thus increase Cn, and Cy,,. A sMall
increase in these quantities was obtained by the addition
of the horizontal tail (figs. 7 and 8). This increment
varied somewhat irreulirly with ai.gle of attack. A
comparison vith the results of reference 5 showed that
the end-plate effect was greaterr for the mic2./ing confieu-
ration. Data from reference 5 in'ricate that this
difference is due to a cn-.rg.e in the v'ing-fusele inter-
ference on the vertical tail with w.n,-, location. In
reference 3 an incremental increase of 0.0010 in Cy, %%as
computed for the end-plate effect of the horizontal tail
on vertical tail L. An average increase of 0.3005 was
obtained from the present ex.---rimental investigation. the
end-plate effect of the horizontal tail on Cli amounted
to less than 10 of effective dihedral. The results of
the present investigation (fig. 8) indicate that, althou-h
separation begins to occur on the vertical tail at about
the same time with the horizontal tail on and off, it
progresses more rapidly with the "-orizontal tail on.

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


Effect of Changes in Fuselege Length

.?. .thin the scope of the present investigation, a
slight increase in Cn,. :as ,r-nerally obtained with a
longer fuselae for the model havi,2' no vertical tail
(fi,js. 9 to 11). T'he effect was more pronounced at the
larger negative an.les of attack.

z?-r the complete model equipped with vertical tail 4,
the increase in CnV with fuselage length was approxi-
mately linear (fi-s- 12 and 15). Tiis increment of Cn:,
which resulted frnom increased fuselage length, appreciably
diminislhid with a positive increase in unele of attack.
This decrease may be no.rtly caused by interference. A
c'.rlp1rison with the results of reference 5 showed that,









NACA ARR No. L5Cl5a


for the midwing configuration, the increase of Cn% v.ith
fuselage length was also linear tut remained fairly
constant with a change in angle of attack.

The variations of 1, and Cy were small, for
the fuselage lengths tested, both witI. and without a
vertical tail. A similar result was ottained for the
midwing configuration (reference 3).


Effect of Changes in Vertical-Tail Area

The slopes Cn i and C ,, increased approximately
linearly with vertical-tail area (figs. 1i- to 17). The
rate of change of Cn, with vertical-tail area was
greatest in a small region tetween an-les of attack of -)jo
and 00 and decreased as the jungle of attack varied from
this range. This change in vartictl-tail effectiveness
with angle of attack. might be attributed to interference.
For the midwing corL'iguration, the increases in these
slopes with vertical-tail area were also approximately
linear and fairly constant over the unstAlled angle-of-
attack range (reference 5).

As would be expected, at negative and small positive
angles of attack, CZv increased with vertical-tail
area whereas, at large positive angles of attack, the
ooposite was true. A similar result was obtained for
the midwing configuration (reference 5).


Effect of Changes with Constant Tail Volume

In figures 13 and 19 the result of changing the
fuselage length and vertical-tail area in such a manner
as to hold the tail volume c instant is shown. The
configurations in which the Lail volume remained constant
are shown in table V. D-ta from figures 13 and 19 are
cross-plotted in figure 1+. The vertical tails tested
all had an asoect ratio of 2.15.
The slone Cn,, should remain anproximately the
same with constant tail volume. The experimental
variation is small over the normal flight range and may
be partly caused by interference or might be explained
by the arbitrary manner in which the tail-volume coef-
ficient was defined.









NACA ARR No. L5C15a


The values of CL and Cy are dependent mainly on tail
area and are practically independent of tail length
(fig. 14). For the range of variations giving constant
tail volume, the changes in both Cy. and Cj,, were
appreciable.


Effects of Changes in Dihedral

For the model having no vertical tail, the change
in Cn* with dihedral angle was small (figs. 9 to 11).
With the vertical tail on, Cny was slightly larger
for F = 00 than for P = 5 (figs. 12 to 17). Similar
trends were also obtained for the midwing configuration
(reference 5). Figure 14 shows 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. The
slope Cy| was generally slightly greater for P = 0
than for P = 5 except at large positive angles of
attack.

The changes with dihedral angle of wing-fuselage
interference and wing-fuselage interference on the
vertical tail were 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 4. The test Reynolds number, based on the
wing chord, was about 619,000 for the Langley 7- by
10-foot tunnel compared with about ..18,000 for the
Langley stability tunnel. The effective Reynolds number,
however, was about the same since the turbulence factor
for the Langley 7- by 10-foot tunnel is 1.6, compared
with less than 1.1 for the Langley stability tunnel.
Data taken from reference 4 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 20 shows









NACA ARR No. L5Cl5a


that satisfactory agreement was obtained, in general, for
all three static-lateral-stability slopes. In both
tunnels the stall occurred at about the sar.e angle of
attack and the model, when yawed, tended to rill violently
at the stall.


CONCLUSIONS


The results of tests of a model consisting of a
rectangular low wing on a circular fuselage with
variations in vertical-tail area and fuselage length
with and without a horizontal tail indicated, for the
range of configurations tested, the following conclusions:

1. The influence of wing-fuselage interference on
the slope of the curve of yawing-moment coefficient
against angle of yaw Cnw and the slope of the curve of
lateral-force coefficient against aa.le of yaw Cyo was
usually stabilizing, apprecisile, and varied with angle
of attack. The effect of wing-fuselage interference on
the values of Cnu and Cy. contributed by the vertical
tail was also generally stabilizing and uporcciable at
negative and small positive angles of attack but varied
with angle of attack.

2. The end-plate effect of the horizontal tail
slightly increased the efficiency of the vertical tail.
The experimental increment obtained was only one-half
the computed value.

5. Increasing the fuselage length with no vertical
tail resulted in a slight increase in Cn, for the
model, both with and without a horizontal tail. At the
larger negative angles of attack, the effect was more
pronounced. For the complete model, the increase in Cngj
was approximately linear with fuselage length. The
magnitude of the increase appreciably diminished with a
positive increase in angle of attack. The changes in
the slope of the curve of rolling-moment coefficient
against yaw C1, and in Cy, with fuselage length were
small.









NACA ARR ITo. L5C13a


4. The slopes Cns, and Cy, increased approxi-
mately linearly with vertical-tail area. For the system
of axes used, CL, increased with vertical-tail area at
negative and small positive angles of attack but the
opposite was true at large positive angles of attack.

5. 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. Except at large positive angles of
attack, CyI was greater with the smaller dihedral angle.


Langley I.iemorial Aeronautical Laboratory
nationall Advisory Committee for Aeronautics
Langley Field, Va.










NACA ARR No. L5ClIa


I'.FL IE;ICEI3


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Qtabtlity in F-wer-Off Flight with Chairts f'or Use
iTn DLsign. NACA hep. No. )b0, 1Qi7.

2. Fehlijcr, Leo F., and l-a:Lachl:v, hob>rt: Investi-
gatio cof .Cffect of Si-'esl!p cn Lateral 7ta".ility
Characteristics. I CIrculur 'se3la;:e wvi.h
Varlatlons n Ver-i.3zal-'ai-il .-.rea r.rL' ?.--il Leia th
with a.-d -th.ut Rcrizc.tal Tii BurfvIe. IACA AAR
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5. Hollingwwar'th, ho as -. : In: T.. ia :n l Lr t .)f
S:iea.Ip on Lte~rA! S1-*ili-r Charactcrl:-tics.
T R- vtcnc' u'lar oI15,i .. j c on Circul.ar .... '.e.laT. v;ith
Variations in VoetS cal-Tail .r c- a '.n 's la,-e
Length v',th R -;J .;Iit'out Hlortzocntal 1ai3l urfaco.
JACA AIL .o. L5C13, l ':..L.

E4. Ea.rber, i.:. J., anr.] Tloi'se, i(. 0. ':Viind-fa-Ji..-l Tnve-Cti-
gation of rEffect of 7;. on Lr-t.eral-- -abilI c
Cl.aracteristics. II .:ect rng-ri.l'r :;.A.C.A. 2 01_
WinE with a Circulr -'.zei ca n 'i .i.,.'A T_,
1!o. 73u, 193:.
5. Viall-ce, A .ithur L."., and Ta.v.,-:-r, ilo ,n' a 1.: 1 ind-
Tunnel InvestLigatin nt Lftfett rf Yaf on Lateral-
Stability Charac teris tics. \ jm.*tr'e .kc l'y
Tapered 'iit w jIi: a Circul.r Fuselaie ilavin,; 'I
Horizontal 'and a Vtrtical -] U.7CA AR1. Jo. 5FJ),
19l5.








NACA ARR No. L5C13a


TABLE I


FTTSELAGE DIMENSIOliS


Fuselage


Short

Loedum

Long


;uselage Tail-cone
length length
(in.) (in.)

2.25 .?.2

37.05 14.65

hil.95 19. 45


Tail length, I
(in.)


20.07

4 37


'Tail length l
Vqinc span b
Wing span b


0 .183
5 lS


29.7 i .6


TABLE II


TAIL-SURFACE DIMEITSIONS


Designation


Vertical-
tail area
(sq in.)
(1)


Vertical-tail area
Wing area


I. I 4 .4


Vertical

Do---

Do---

Do---

Do---

Horizontal


10.835

25.78

28.57

55.16

46.20

64.21


0.0500

.0659

.07.36

.0974

.1230

.173


iArea measured from root chord at


center line of fuselage.


NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS


Tail
surface


aspect
ratio


2.15

2.15

2.15

2.15

2.15

3.99









NACA ARR No. L5C1Ia 17



TABLE TII

rVODEL COMBIUATIODS TESTED


Horizontal Vertical Dihe ral
1l ta. i Auselage I angle Variable
tail tl i

-ff f


2 Short,

Sand L:. n,-

n I -

Onho t











Off L a
O------ --- Long









I Off I 5 nd
O f f-- Long ---------


4 Short 0 and 5 a and



NATIONAL ADVISORY
COI.,T:,ITTEE FOR nEROtJAUTICS









NACA ARR No. L5C15:


TABLE IV


PRESENTATION OF RESULTS


Figure Description of figure 1 Data
presented
3 Lift and drag curves fxr repre- CL and
tentative model confiLurations CD as fi )

4 Slone of yawing-mLnoment and 'Cnl and
lateral-fDrce coefficients for CY,' as f(a)
UiACA 25012 reccnrgular wing :

5 Effect of wing-fuselage itlCn., and
interference A1Cy as f(a)

6 Effect of wing-fuselage I2 Cn, and
interference on vertical tail A2C-',' as f(a)

7 End-plate effect of horizontal Cni, Cj, and
tail CY, as f(u)

3 End-plate effect of horizontal Cn, Cl., and
tail Cy as fis)

9 Effect of changing fuselage lengthlCn C,, and
(no tail surfaces) I Cy, as f(a)

10 Effect of changing fuselage length Cn, Ct, and
(no tail surfaces) I Cy as fII)

11 Effect of changing fuselage length C,1,, Cid,, and
(horizontal tail on; vertical I Cy'u as fla)
tall off)

12 Effect of changing fuselage length Cn,, CZp,1 and
(horizontal tail and vertical CYur us f(u)
tail 4 on)

15 Effect of changing fuselage length Cn, Cl, and
(horizontal tail and vertical I Cy as f(lr)
tail 4 on)


NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS









NACA ARR No. L5Cl5a


TriLE IV Concluded


PRESENTATION OF RESULTS Concluded


Description )f figure Datea
presented


Effect of changing fuselage length



Effect of changing vertical-tail
area

Increment of slope of yaw'ing-
nmment coefficient against
angle of yaw caused by
vertical-tail area

Effect of changing vertical-tail
area

Effect of changes wLth tail volume
constant

Effect of changes with tail volume
constant

Comparison of data from Langley


stability and Langley 7- by
10-foot tunnels


'r,, CL,, and
Cy, as f -:


CnLrp CLI,, and
Cyf as f(a)


nas f -a



Cn, CL, and
Cy as f( )

Cn,i, C01, and
Cy, as f(u)

Cn, CL, and
Cy as f(&)

Cny, Cas,, and
Cy as f(a)


NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS


Figure


14



15


16




17


19


19


20











NACA ARR No. L5C13a ,0






].C




0 C0 0



-4
0 0





P E-
0--4 "
E- I i Hi


-4 I





S.-0d M rj.
W (U V- Go US
S1.1 0 0 0



C,





h, 0 rO 03 O 0


2 r--i




0 00



--4




I rj I
U d rii___ OI








NACA ARR No. L5C13a


TAIL 3
1.0 28 P


T 4 78




TAIL 4


TAIL 5


TAIL 2


A AND 8 ARE QUADRANTJ
OF SIMILAR ELLIPSES.


Figure /.- Rectangular NACA 230/2 w/nin n c wnbination with
circular fuselage, vertical and hormonfal fails, and
/oil cones All dimensions g9ven in inches .


Fig. 1










NACA ARR No. L5C13a Fig. 2












r .










-4



0) -4




bo








o0t

04
o




0
4-+
IC




Ola

.0






14.

4) $4


bo
-.4








NACA ARR No. L5C13a


.4







-


0


-.2
-.A<


-/Z -8 -4 0 4 8 /2 16 20

Angle of offoc*A, cC, o de

Figure 3.- Vao-,'r,on of /,it and drag coel/,cents wi/tl on'/e
of attack for representative model cont,,urations. i; 5
Y, 0.-; 65- /b/sc .(I


Fig. 3






NACA ARR No. L5C13a


IZ



a ^o
MI












4 to
'10
QJ b!


Fig. 4








NACA ARR No. L5C13a


2C

\ p==































0


O


qj








Kj


S0>
L

O


I)
b

'3


4o-


Fig. 5






Fig. 6a NACA ARR No. L5C13a *





:iE
"--Z Z ?a I






--4_ -4- t3












--- -


cz
N






NACA ARR No. L5C13a Fig. 6b














0 + Its%
or






_4 _,*

I,


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NACA ARR No. L5C13a


SI I


























Ll
S \
!& 5 S.
- S
- I "




- ? %'1 k











~~z^I


~bL
N


Fig. 7a








NACA ARR No. L5C13a


' t 'j


'Kj


Fig. 7b







NACA ARR No. L5C13a


.04 --
Horizotao/ Verftca


_-- x Off Off
o Off 4
S___+ On 4
















z
Iq)
k O x



N TiONAL ADVISOR(
COMM TTEE FO AERON ITICS





(a) / 00 *
-/,ure 8. Effect of horizontal/ taol jaurface on
vari0o-on of yawing /f--moment rol//g -
moment and /ateral force coeff1cientn with
ong e of y w Jhort fufse/age ; a l0.2


Fig. 8a







NACA ARR No. L5C13a


.02












./0








- 0
JeO


Hor z on to/ Vert/cal
tfo/ tall
Off Off
o Off 4
+ On 4


o I
-,%
'Il
^IZ.I


-30 -zO -/0 0


Angle of yaw ,


-W de g


fb) F/ 5 o.
Figure 8 Concluded .


Fig. 8b








NACA ARR No. L5C13a


t)
0


N


















CO
0
OSo
' 0
1.*
OO
^ $'S




U &)
^ ^-



qj b~ <
Q, I~t

^S: 0 a


AL


Fig. 9







NACA ARR No. L5C13a


.02

Z
at'
I~0




















ttj
II'

it
%.I)

t'tha/0
k


-30


-2 0 -/0 0 /0


S**k
'o

0 *<



It
Qc


Angle of ya w


?/gure /O Effect of changing fuselage /length
on variation of yowing moment ro/llin -
moment and lateral/ force coeffltenAr
with ang/e of yaow //orizona/l anod verfico/
fo/al off; aI /i0. 02 I / 5 ; q 65 lb/sq ft.


Fig. 10






Fig. ha NACA ARR No. L5C13a





-F _--^- T~ __
-4,

ozo





.r a


lb


iZ Z
\ O 4 O
^ "'-V II--r .
-- 3 --I-- ^ L ------ __ z .
t o ^ x '_ < I..'b



k,
----_ 4---- sB t^ t


i -----^ ^ -4- I






0
____ _____ K_[o


co






NACA ARR No. L5C13a


Fig. lib




N


ccc



Sj


'a
'3


'3







Fig. 12a


NACA ARR No. L5Cl3a




I/



0 C


















Q)
cz
0 uN
-^ -

*i? OO

^_O 0l


~I.






NACA ARR No. L5C13a


41



|
ii
1s

0





I 4-


q I


~I. A


Fig. 12b







NACA ARR No. L5C13a


.04



.02
o"






-.02









tb
, *0
N,-.02


qJ


-30 -20 -/0 0 /0
Ang/e of yaw 1~ dog

fo r o 0
/gure I/ Effect of fuse/lgye /ength on
var/at/on of yow/n; moment ro///ng moment
and /ateral force coeff/c/entf w1dh ong/e of
yaw /or/zon ta/ toll and vert/ca/ faol 4 on. a, /O.
a, 65 fb/s g ft .


Fig. 13a







NACA ARR No. L5C13a


.04


t02





qj

-.02








.10


C'
'K
~-d0
'a

'3
0


-30


FiLse/ape


Short"
melll


'K




ciz

13
N

-20 -/0 0 /0


Angle of yaw


, j ,de g


(b) /F, y .
Fif-gre / Concluded .


Fig. 13b






NACA ARR No. L5C13a


D//hedro/
(deg)
-----


(de~,7)


0


C,7


.0021--


-.004



.002



0

.0/

67


N


o Short
+ meAa/Um7
x Lo7g
o S/?or/
0> AMd'/7um
A Long


t7/ vo/ me, 0.0407 _












T- IONAL ADVISOR Y
-4 aCMM ITIEE F AERON UTICS


.04


.08


./ZI


./6


(a) c ,0 0
Figure 14.-Effect of chranygnq fuse/age
length on var/ition of /ateroa-stlahi/y
slopes C,~), C ,, an CyV with vertcal/-
tail area. Hor/'zonta/ toil on
7, 65 lb/sq f t.


Fig. 14&a







NACA ARR No. L5C13a


Dhedro/
(deg)
----5


---- 0


0


-.00Z



-.004

.00,e


C2*


cyi


Fuae/lage


Dlheo'o/


0
0
0
5
5
5
if


0 .04


.08


J'v/'Sw
(bt) a,/0.


F/gure /4 .- Concluded.


./2


.1/6


Fig. 14b






Fig. 15a NACA ARR No. L5C13a


.


o ok














IIC)
a4o
-fe-t tH -^-^ ^^,lit,^


-L^-^~~~~ [1h ---^ ^


1~~~~~~c q --,t ~l '*







NACA ARR No. L5C13a


+ x 0o <
r 4


---- -t
-

) x -
\u/

.



o5Sto


c '







-


I
O


Fig. 15b


o
b


re
6







NACA ARR No. L5C13a


0-




-00-0--


-.006
NATKC NAL AD ISORY
CMM__ITfE FOR A RONAUTIS



0 .04 ./0 /6



Figure /6.- Increment of /laere/-st sb////
slope ACn,7 caused by vertical- tail area.
(C7. for mnedlum-/ength fuse/oA e w/h horizontal
foil ond wing with 0 dihedral subtracted from C,-
for complete model to ob a/n A Cn,.) q,65 l/b/rq fr.


Fig. 16







NACA ARR No. L5C13a


.04







| 0 ---- Ver/cal


0 3
-.02- o 4 \ .02







./0







GO N, TIONAL ADVISOI
4 COMMITI TEE FO AERON, ITICS


-30 -0o -/0 0 /0
Angle of yaw 36- de,

(a) /-oq0
Figure 17. Effect of chag/n ver/cal ta/
area on variaton of yawing moment rollin -
moment and lateral-force coefflclent.x with
angle of yow AMedom-/en-gth fuselage with
horizont/al tail ; a /0.2 ; q 65 lb/lg ft.


Fig. 17a








NACA ARR No. L5C13a


- ~'-- p y y I


t4'r/Vca/
fai

0 3
4


6- 0-02



0












./0
4z


I~3


N OIONAL ADVISOR
COMM TTEE FO AERONI UTICS


-/0


Angle of yaw < deg

(b) /' 5.
Figure /7 Concluded .


__ o2


~ ~ ~ ~ ~7 L


0




-./0

-)
-,0


-30


-20


Fig. 17b


*w *"-







NACA ARR No. L5C13a


-oo
S I0















-4,
0%0
v j Q .









b




*- L








'00

0'


A
1~~
(3 C~)


Fig. 18a







NACA ARR No. L5C13a


-0 a




8
-o-
1.* '-






B
Qp

!



-


iz
Q


'Fig. 18b







NACA ARR No. L5C13a


.04







d









10



1w
Is'


-30


-20
A nqle of


-/0
yaw 5


Fuase'/aye


o Short
x A4edum
SLot,;


0
) 5 deg


(a) / 0 .
Figure /9 Effect of revera/ comblnotions havn;/
con-ontf tail volume on vor/ation of yawing -
moment rol//ing-moment, and loferalo/-force
coeffi/cents w/th ongle of yaw Tail volume y
0.0407; hor'zo nfa/ tai/ on; a /0.2; q 65/b/rg ft.


--- --- ---- ---- --- ------1 ,Jp'-- --- --


____ ____4____ ____________

y, NA IONAL / ADVISORY
~_____ ~___ --- ---COMMIT EE FOR ERONAUI


____________^?,,____
Y
--------Y-
Xs.
__________________________________^s__
X




__^^^^_-T^_


-1 i


Fig. 19a


Vertila/
tadl


3







-.02







NACA ARR No. L5Ci3a


I


















>0


. -
a o;


A/7gle of


Vert/ca/
ta/l










4 '


.oe
-^ 4'


-0


-/O 0


yaw


,

(b) r/ 5.
Figure /9 Conc/uded .


Fig. 19b


-O -
-130 -12







NACA ARR No. L5C13a


o' o


-Q
tl-


r G

Q),
LZ



0 0c:
'k












Lz,
qj N, -,
o 4







OS
LO


* -7


jbL
(~.)


Fig. 20











UNIVERSITY OF FLORIDA


3 1262 08104 942 0





-. ..L S DEPART.'vENT
2 i MiRSTON SCIENCE LIOWARY
i(U. 0BX 117011
AIlNESVILLE,. FL 321 ?-70Q t USA



UNIVERSITY OF FLORIDA
DOCUMN---TS D-, .' '.T

- ....,
I. 7 -;,'1-


:i L I i. -', 'Y. -