I
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I
SZ 70131
NASA MEMO 1-15- 59L
COPY
NASA
MEMORANDUM
EFFECTS OF FUSELAGE NOSE LENGTH AND A CANOPY ON THE
LOW-SPEED OSCILLATORY YAWING DERIVATIVES OF A
SWEPT-WING AIRI°LANE MODEL WITH A FUSELAGE
OF CIRCULAR CROSS SECTION
By James L. Williams and Joseph i%. DiCamillo
Langley Research Center
Langley Field, Va.
NATIONAL AERONAUTICS ANDSPACE ADMINISTRATION
WASHINGTON
January 1959
NATIONAL AERONAUTICS AND SPACE ADMINISTRATION
MEMORANDUM 1-15-59L
EFFECTS OF FUSELAGE NOSE LENGTH AND A CANOPY ON THE
LOW-SPEED OSCILLATORY YAWING DERIVATIVES OF A
SWEPT-WING AIRPLANE MODEL WITH A FUSELAGE
OF CIRCULAR CROSS SECTION
By James L. Williams and Joseph R. DiCamillo
S_Y
A wind-tunnel investigation was made at low speed in the Langley
stability tunnel in order to determine the effects of fuselage nose
length and a canopy on the oscillatory yawing derivatives of a complete
swept-wing model configuration. The changes in nose length caused the
fuselage fineness ratio to vary from 6.67 to 9.18. Data were obtained
at various frequencies and amplitudes for angles of attack from 0° to
about 32 ° . Static lateral and longitudinal stability data are also
presented.
INTRODUCTION
The results of previous wind-tunnel investigations (refs. I to 4)
have indicated that wings of swept design have lateral oscillatory sta-
bility derivatives that become increasingly large at high angles of
attack. These results also showed that the oscillatory derivatives are,
in some cases, substantially different from the steady-state derivatives.
Some results of reference 3 have shown that the large magnitude of the
derivatives at high angles of attack is dependent to some degree on
frequency and amplitude of the oscillatory motion. There are certain
airplane parameters, also_ which may have a modifying effect on the
magnitude of these oscillatory derivatives. For fuselages with square
cross section, for instance_ the fuselage nose length and the canopy have
considerable effect on certain static stability derivatives. (See
ref. 5-) No data on the dynamic derivatives were given in reference 5,
however.
In the present investigation the oscillatory technique of reference 3
was employed for the purpose of determining the effects of fuselage nose
length and a canopy on the oscillatory lateral stability derivatives of a
2
complete swept-wing model with circular fuselage cross section at vari-ous frequencies and amplitudes.
COEFFICIENTSANDS_q4BOLS
The data are presented in the form of coefficients of forces andmomentswhich are referred to the system of stability axes with theorigin at the projection on the plane of symmetry of the quarter-chordpoint of the meanaerodynamic chord. The oositive directions of forces,moments, and angular displacements are shownin figure i. The coeffi-cients and symbols used are defined as follows:
b
!
CD
wing span, ft
approximate drag coefficient,Aoproximate drag
qS
LiftCL lift coefficient,
qS
C_ rolling-moment coefficient, RolLing momentqSb
Cm
pitching-moment coefficient,Pitching moment
qS_
C n
c
k
q
r
yawing-moment coefficient, Yawi.%g moment
qSb
wing chord, ft
wing mean aerodynamic chord, ft
reduced frequency parameter, mb/2V
1 2 ib/sq ftdynamic pressure, _V ,
angular velocity in yaw (r = _), radians/sec
_r radians/sec 2= _-_,
3
wing area, sq ft
t
V
time, sec
free-stream velocity, ft/sec
c_ angle of attack, deg
angle of sideslip, radians or deg
$_ radians/sec
P mass density of air, slugs/cu ft
angle of yaw, radians or deg
_o
co
amplitude of yawing oscillation, deg
circular frequency of oscillation, radians/sec
(_C_ _n
C_r - _. Cnr - _b_
2V 2V
8C t 6C_ Cn" _ n
CZ_ _ r $}b 2
4V2 4V 2
_C_ _C n
CZ_ = _7" Cn_ - _
8c_ : 8C___n_n
Subscript:
40, 45, 50, 55 overall fuselage length, in./
The subscript _ when used with a derivative Ifor example,
C_,_ + k2C_r,m.) indicates that the derivative was obtained from an
oscillation test.
4
Model designations:
F
W
VH
WF
fuselage
wing
vertical and horizontal tails
wing and fuselage
APPARATUS
The apparatus used in the present investigation for the oscillation-
in-yaw tests is described in detail in reference 3. The oscillatory
rolling and yawing moments were measured by a two-component resistance-
type strain gage attached at the assumed center-of-gravity location of
the models. The output signals from the strain gage were modified by a
sine-cosine resolver so that the measured signals were proportional to
the in-phase and out-of-phase components of the strain-gage signals.
These signals were read on a highly damped direct-current meter. This
recording equipment is described in detail in the appendix of reference 1.
MODELS
Drawings of the models used in the present investigation are pre-
sented as figure 2, and a photograph of a model is presented as figure 3.
Pertinent geometric details are given in table I. In order to maintain
about the same amount of directional stability for each model at _ = 0°,
a different size vertical tail (with aspec_ ratio of 1.4) was used with
each fuselage. All model components (wing, fuselage, and tails) were
made of balsa wood with a fiber-glass covei:ing. The wing and tail sur-
faces had a 45 ° sweptback quarter-chord line, a taper ratio of 0.6, and
NACA 65A008 airfoil sections parallel to the airstream. The wing and
horizontal tail, which were common to all models, had aspect ratios of
3 and 4, respectively, and each was mounted in a low position on the
fuselage. The fuselages were of circular ._ross section with a pointed
nose and blunt trailing edge. The fuselag,_ fineness ratio varied from
6.67 to 9.18. (Fuselage length varied from 40 inches to 55 inches.)
Fuselage coordinates are given in table II. The canopy dimensions
selected were average values determined frc_m several present-day fighter-
type airplanes. The canopy was located at the same distance from the
nose of each fuselage, and thus its distan,_e from the tall assembly
varied with the length of the fuselage nos_. (See fig. 2(b).) Canopy
coordinates are given in table III.
TESTS
All tests were madein the 6- by 6-foot test section of the Langleystability tunnel (ref. 6) at a dynamic pressure of 24.9 pounds persquare foot, which corresponds to a Machnumberof 0.13. The testReynolds numberbased on the meanaerodynamic chord was approximately0.83 x 106. The oscillation tests consisted of measurementsof the in-phase and out-of-phase rolling and yawing momentsfor a range offrequencies and amplitudes. The WF5oVHconfiguration was oscillated atfrequencies of 0.5, 1.0, 1.5, and 2.0 cycles per second at amplitudes ofyawing oscillation of T2° , ±6° , ±i0 °. These frequencies correspond tovalues of the reduced-frequency parameter _b/2V of 0.0282, 0.0564,0.0846, and 0.1129. Breakdowntests were madeonly w_th the WF50VHcon-figuration at 1.5 cycles per second and an amplitude of yawing oscillationof _6°. The effect of a canopy on the complete model configurations forthe various fuselage lengths was also determined only at a frequency of1.5 cycles per second and an amplitude of yawing oscillation of T6° .
For each amplitude, frequency, and angle-of-attack condition, awind-on and a wind-off test was made. The effects of the inertia of themodel were eliminated from the data by subtracting the wind-off resultsfrom the wind-on results.
The static derivatives C_ and Cn_ were obtained from tests at= 0° and _ = ±5° with the sameequipment that was used for the
oscillation tests. The lift_ drag_ and pltchlng-moment results weremeasured (at _ = 0°) by means of a six-component mechanical balancesystem.
For all tests, oscillatory and static_ the angle of attack rangedfrom 0° to about 32° .
CORRECTIONS
Approximate jet-boundary corrections as determined by the methodof reference 7 were applied to the angle of attack and the drag coeffi-cient. For the configurations with horizontal tail, the pitching momentwas corrected for the effects of Jet boundary by the methods of refer-ence 8. No Jet-boundary corrections were applied to the oscillatoryresults.
The data are not corrected for the effects of blockage and support-strut interference.
6
RESULTS
The results of the investigation are presented in the following figures:
Figure Coefficients plotted _o, _bagainst _ deg 2-V Configurations Canopy
C n - Cn .
Cn_,_ + k2Cn_,_
C_r,m - C_,m
C_ + k2Cz.
Cnr,_ - Cn_,m
Cn_,_ + k2Cn_,_
C_r,_ - C_,_
C_,_ + k2Cz_,_
Cnr, _ - Cn_,_
Cn_,_ + k2Cn_,_
C_r,_ - C_,_
C_,_ + k2C_.r,_
Cn_, CI_
Cnr, _ - Cn_,_
Cn_,_ + k2Cn_,_
C_r, _ - CZ_,_
C%_,_ + k2Cz_,_
l
C m, CL, C D
+-6
t2, _6, ±i0
±2, Z6, ±i0
+_6
0.0846
O.0282
.0964
.0846
.1129
O. 0282
•0564
•0846
•n29
o.o846
WF49VH
WFsoVH
WF55W
WFsoW
WFsoW
WF49W
_0 _
WF_sW
WFso
_O _
WFsoVH
_4o_
WF45W
_o _
WF55W
On and off
Off
Off
On and off
Off
On and off
7
Increasing the fuselage nose length by as much as 75 percent and
making compensating increases in tail size did not have an undesirable
influence on the variation of yaw damping and directional stability
with angle of attack. Substantial influences of canopy addition were
apparent_ however (fig. 4(a)). The effects of changes in frequency and
amplitude of motion were also significant. Such effects have been
noted in reference 3 for wings alone, but not to such an extent at the
lower angles of attack as is shown in the present results for changes
in amplitude (fig. 6(a)).
Langley Research Center,
National Aeronautics and Space Administration,
Langley Field, Va., October i, 1958.
8
REFERENCES
i. Queijo, M. J., Fletcher, Herman S., Marple, C. G., and Hughes, F. M.:
Preliminary Measurements of the Aerodynamic Yawing Derivatives of a
Triangular, a Swept, and an Unswept Wing Performing Pure Yawing
Oscillations, With a Description of the Instrumentation Employed.
NACA RM L55LI4, 1956.
2. Campbell, John P., Johnson, Joseph L., Jr., and Hewes, Donald E.:
Low-Speed Study of the Effect of Frequency on the Stability Deriva-
tives of Wings Oscillating in Yaw With Particular Reference to
High Angle-of-Attack Conditions. NACA RM L55H05, 1955.
3- Fisher, Lewis R.: Experimental Determir_tion of the Effects of
Frequency and Amplitude on the Latera] Stability Derivatives for a
Delta, a Swept, and an Unswept Wing O_cillating in Yaw. NACA
Rep. 1357, 1958. (Supersedes NACARM L56AI9.)
4. Riley, Donald R., Bird, J. D., and Fisher, Lewis R.: Experimental
Determination of the Aerodynamic Derivatives Arising From Accelera-
tion in Sideslip for a Triangular, a Swept, and an Unswept Wing.
NACA RM L55A07, 1955.
5. Jaquet, Byron M., and Fletcher, H. S.: Effects of Fuselage Nose
Length and a Canopy on the Static Longitudinal and Lateral Stability
Characteristics of 45 ° Sweptback Airp]ane Models Having Fuselages
With Square Cross Sections. NACA TN _961, 1957.
6. Bird, John D., Jaquet, Byron M., and Co, an, John W.: Effect of Fuse-
lage and Tail Surfaces on Low-Speed Yawing Characteristics of a
Swept-Wing Model As Determined in Cur_ed-Flow Test Section of
Langley Stability Tunnel. NACA TN 24_3, 1951. (Supersedes NACA
L8GI3.)
7. Silverstein, Abe, and White, James A.: Wind-Tunnel Interference With
Particular Reference to 0ff-Center PoEitions of the Wing and to the
Downwash at the Tail. NACA Rep. 547, 1936.
8. Gillis, Clarence L., Polhamus, Edward C., and Gray, Joseph L., Jr.:
Charts for Determining Jet-Boundary Corrections for Complete Models
in the 7- by 10-Foot Closed Rectangulsr Wind Tunnels. NACAWR L-123,
1945. (Formerly NACA ARR L5G31.)
9
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TABLE II.- FUSELAGE COORDINATES
x_ in.
024
68
io12141618
2o2224
2628
5o525456
584O
4244
4546
48
5O5254
55
R40 , in.
0.64
i .201.68
2.092.42
2.672.852.963 .oo2.972.902.80
2.68
2.552.402.262.10
1.92
Z.721.50
R45 , in.
0.64
i .201.68
2.092.42
2.672.85
2.965.oo2.992.972.95
2.872.792.7o2.60
2.472.552.3_82 .Ol1.82i .61
1.50
R_IO_ in. R55 , in.
i
i222
22
5
35.005.002.992.952.902.852.752.652.542.402.262.10
1.92
i.72i.50
0• 64 .64.20 i.20.68 1.68
.09 2.09
.42 2.42
.67 2,67•85 2.85.96 2.96
.00 3.00•O0 3.O0
3.00
5.oo5. oo5.oo2.992.972.95
2.872.792.702.602.47
2.352.182 .Ol1.821.61
1.50
ii
TABLE III.- CANOPY COORDINATES
14. O0 in.Z
x_ y_ z_ R_ x_ y_ z_ R_
in. in. in. in. in. in. in. in.
i
2
3
0 1.68 1.68 1.28 2.O64
o 3.840.70 1.75.64 1.90
•55 2.o5.44 2.20
.29 2.350 2.47
0.97 1.85
0 3.13
1.16 1.98
1.o4 2.25
•93 2.50
.80 2.75
.66 3.00
.51 3.25
.27 3.50
o 3.60
1.895
2.09 6
2.28 7
8
1.30 2.180 4.00
1.3o 2.351.2o 2.60
1.o7 2.85
•95 3.1o
.81 3.35
.67 3.60
.46 3.85
o 4.09
1.23 2.49
0 4.00
2.42
2.55
2.67
2.77
2.85
xj y_ z_ R,
in. in. in. in.
0.95 2.75
.84 3.OO
.68 3.25 2.909 .50 3.50
.22 3.75
0 3.80
lO 0.78 2.85 2.96o 3.64
0.59 2.93 2.99II0 3.50
0.40 2.98
.34 3.06
12 .25 3.16 5.00
.14 3.26
0 3.36
13 0.19 2.99 3.00
14 o 3.00 3.00
12
Y
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X owing mo
I_ _ i L .... _ _ _ " _
_zontol ref. pIQne
Rolhng moment
×_.l
F_elalwe wind
Lifl
Pitchmg m3ment _ Rolling moment
C#
z
Figure i.- System of stability axes. Arrows indicate positive direc-tions of forces_ moments_ and angular displacements.
13
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Figure 2.- Conclud(_d.
15
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£±gu_e _.- £££ect o£ _e_ce& £_e_Qency p_&_ete_ o_ st&biZ±ty _e£±va-
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19
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0 4 8 12 16 20 24 28 32 0 4 8 12 16 20 24 28 32
Angle of ottock, OS,deg Angle of ottock, 0_, deg
(b) ¢o = t6o.
Figure 5.- Continued.
2O
12
8
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=
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20
112
=12
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0 4 8 12 16 20 24 28 32
Angle of o/rock, CC, deg
J
0 4
(c) _o = +lO°"
Figure 5.- Concluded.
8 12 16 20 24 28 32
Angle of attock, OC,deg
21
22
2O
/6
/2
04
0
-04
0 4
(b) Directional-stability characteristics.
Figure 6.- Continu_:d.
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