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MEMORANDUM 


for the 

U. S. Air Force 

INVESTIGATION IN THE LANGLEY FREE-FLIGHT TUNNEL OF THE 

LOW-SPEED STABILITY AITD CONTROL CHARACTERISTICS 

OF A l/lO-SCALF MODEL SIMULATING THE 

CONVAIR F-102A AIRPLATTE 

By Peter C. Boisseau 

Lsingley Aeronautical Laboratory 
Langley Field, Va. 


class; 

This material contains information affecting^ 
of the espionage laws, Title 18, U.S.C.i 
manner to an unauthorized person is prohil 


DOCUMENT 



Defense of the United States within the meaning 
, the transmission or revelation of which in any 


NATIONAL ADVISORY COMMITTEE 
FOR AERONAUTICS 

WASHINGTON 


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MCA EM SL55B21 


^^CLASSIP 

SKiPHlie 

MTIOML ADVISORY COMMITTEE FOR AEROMUTICS 




RESEARCH MEMORANDUM 
for the 

U. S. Air Force 


INVESTIGATION IN THE LANGLEY FREE-FLEGHT TUNNEL OF THE 
LOW-SPEED STABILITY AND CONTROL CHARACTERISTICS 
OF A 1/lO-SCALE MODEL SIMULATING THE 
CONVAIR F-102A AIRPLANE 
By Peter C. Boisseau 

SUMMARY 


An investigation of the low-speed, power-off stahiliiy and control 
characteristics of a l/lO-scale model simulating the Convair F-102A air^ 
plane has heen made in the Langley free-flight tunnel. The model in 
its hasic configuration and with tvro modifications involving leading- 
edge slats and an increase in vertical-tail size was floim through a 
lift-coefficient range from O .7 to the stall. Only relatively low- 
altitude conditions were simulated. 

The longitudinal stability characteristics of the model were con- 
sidered satisfactory for all conditions investigated. The lateral 
stability characteristics were considered satisfactory for the basic 
configuration over the lift-coefficient range investigated, except near 
the stall, where large values of static directional instability caused 
the model to be directionally divergent. An 80-percent increase in 
vertical-tail area increased the angle of attack at which the model 
became directionally divergent. The longitudinal and lateral control 
characteristics were generally satisfactory. Although the adverse 
sideslip characteristics for the model were considered acceptable over 
the angle-of -attack range, analysis indicates that the adverse sideslip 
characteristics of the airplane may be objectionable at high angles of 
attack. 


2 


MCA EM SL55B21 


IMTRODUCTION 


An investigation of the low-speed stability and control character- 
istics of a l/lO-scale model simulating the Convair F-102A airplane has 
been made in the langley free-fli^t tunnel at the request of the 
U. S. Air Force. !The F-102A airplane is a turbojet-powered^ interceptor- 
type airplane irith a 6o° delta i^ring and a 6o° delta vertical tail. It 
differs from the Convair TF-102 airplane of reference 1 by having a 
longer fuselage, a longer tail moment arm, a drooped leading edge which 
increases the iri.ng area slightly, and chordwise fences at the 65 percent 
■(•ring semispan station. These changes were made to the free-flight- 
tunnel model of the IF-102 from specifications furnished by Convair in 
September 1955* Any changes to the full-scale airplane subsequent to 
this date were not incorporated in the free-flight-tunnel model. 

The investigation included fli^t tests of the model in its basic 
configuration and ^rLth several modifications involving leading-edge 
slats and an increase in vertical- tail size. Force tests of these con- 
figurations were also made to determine the static stability 
characteristics . 

In order to permit a better interpretation of the free-flight- 
tunnel tests in terms of the full-scale airplane, a conqjarison was made 
betoeen the results of force tests at a low Reynolds nxjmber (O .85 x 10^) 
in the free-flight tunnel and force tests made by Convair at a higher 
Reynolds number ( 5 . 316 x I 06 ) . 


SYMBOIS 


All stability parameters and coefficients are referred to the 
stability system of axes originating at the center of gravity. A sketch 
sho^ang the axes and the positive directions of the forces, moments, and 
angles is given in figure 1 . 

S wing area, sq ft 

S-j; exposed vertical-tail area, sq ft 

c mean aerodynamic chord, ft 

V airspeed, ft/sec 

b 


<1 


wing span, ft 

dynamic pressure, Ib/sq ft 



•• 


MCA RM SL55B21 


3 


• • 

!••• 

• • 

fO 9 

99 


P 

\J 

m 

P 

t 

01 


n 

ix 

ly 

iz 

% 

% 

X 

Y 

Z 

M 

H 

L 


air density, slugs/cu ft 
v/ei^t, Ib 
airplane mass, slugs 
relative-density factor, m/ pSb 
angle of sideslip, deg 

angle of yaw, relative to the tunnel axis system, deg 

for flight-test data, the angle hefereen the projection of 
the Y-axis of the model on the YZ-plane of the tunnel 
and the Y-axis of the timnel, deg 

angle of attack, deg 

inclination of principal longitudinal axis of airplane trith 
respect to flight path, positive when principal axis is 
above fli^t path at the nose, deg 

moment of inertia about longitudinal body axis, mkj^, 
slug-ft^ 

moment of inertia about lateral body axis, mky^, slug-ft^ 

moment of inertia about normal body axis, mkg^, slug-ft^ 

radius of gyration about longitudinal body axis, ft 

radius of gyration about lateral body axis, ft 

radius of gyration about normal body axis, ft 

longitudinal force, lb 

lateral force, lb 

normal force, lb 

pitching moment, Ib-ft 

ya-vri.ng moment, Ib-ft 

rolling moment, Ib-ft 



k 


MCA EM SL55B21 


Cl 

Cd 

Cm 

Cn 

Cz 

Cj 


lift coefficient, lift/ q.S 
drag coefficient. Drag/ qS 
pitching-moment coefficient, M/qSc 
yawing-moment coefficient, IJ/ qSb 
rolling-moment coefficient, L/qSb 
lateral-force coefficient, Y/qS 


s°lr 


Sr 


rudder deflection, deg 


5g elevator deflection (elevons deflected together for 

elevator control) , deg 

Sg, aileron deflection (elevons deflected differentially for 

aileron control) , deg 


APPARATUS AND MODEL 


The flight tests and static force tests were conducted in the 
langley free-flight tunnel, which is designed to test free-flying 
dynamic models. A complete description of the tunnel and its operation 
is presented in reference 2. Force tests were made ’irith a sting- type 
support system and an internally mounted strain-gage balance. 

The l/lO-scale model used in the investigation was obtained by 
modifying the original Convair YF-102 model used in the investigation 
of reference 1 so that it approximated the fuselage shape and accurately 
represented the other geometrical changes of the revised design, such 
as fuselage length, vertical-tail position, leading-edge mng droop, 
and wing fences. A three-view draT'fing of the model is shoim in fig- 
ure 2 and a photograph of the model is sho™ in figure 5* Table I gives 


MCA RM SL55B21 


5 


the scaled-up mass and dimensional characteristics of the model. 

MLdspan leading-edge slats and two different sizes of vertical tails 

were also tested on the model. (See fig. 2.) !Ehe vertical tails 

. 

tested were the basic tail I — & = 


increase in area 


( 1 = 0 . 18 ), 


if = 


and a tail with an 80 -percent 


DEOERMIWATION OF STATIC STABIIZTT AND CONTROL CHARACTERISTICS 

OF FLIGHT-TEST MODEL 

Force Tests To Determine Longitudinal Stability and Control 


Force tests were made to determine the static longitudinal stability 
and control characteristics of the basic model and the model with modi- 
fications for an angle-of-attack range from 0° throu^ the stall. All the 
force tests were run at a dynamic pressure of 5*63 pounds per square foot, 
which corresponds to an airspeed of about 55*7 feet per second at standard 
sea- level conditions and to a test Reynolds nmber of O .85 X 10^ based on 
the mean aerodynamic chord of 2.32 feet. 

Static longitudinal characteristics of the basic and modified model 
are presented in figure 4. The data are presented for a center of 
gravity of 30*0 percent of the mean aerodynamic chord in order that 
comparisons can conveniently be made with the Convair IF-102 data from 
reference 1. The leading-edge slats of figure 2 were used on the model 
because they had a beneficial effect on the lateral stability character- 
istics at high angles of attack for the original model of the 
Convair yF-102 tested in the langley free-flight tunnel (ref. l) . The 
data of figure 4 indicate that these slats were obviously not the optimum 
configinration for producing the most satisfactory longitudinal character- 
istics for the model of the present investigation. The data show that 
the slats decreased the lift-curve slope, the maximum lift coefficient, 

and the static longitudinal stability parameter ^ comparison of 

the pitching-moment curves for the t^-ro conditions shows that the slats 
caused a slight pitch-up at the stall. 

The static longitudinal stability and control characteristics of 
the free-flight-tunnel models of the F-102A and IF-102 are presented in 
figure 5- These data show that the lift-curve slopes and the maximum 
lift coefficient for the F-102A are greater than those of the IF-102. 

The increase in the maximum lift of the F-102A over that of the IF-102 
can be attributed mainly to the cambered leading edge of the F-102A. 



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MCA EM SL55B21 


About 25 percent of the increase can be attributed to the fact that the 
1***^ coefficients vrere based on the area of the YF-102 \ring, which was 

; approximately 5 percent less than that of the F-102A wing. A comparison 

•••• of the pitching-moment curves shows that the models had about the same 

® static longitudinal stability and elevator effectiveness over the lift- 

coefficient range. 

• •• 


Force Tests To Determine lateral Stability and Control 

Force tests were made to determine the static lateral stability and 
control characteristics of the model >rith the vertical tail off and on 
over a sideslip range from 20° to -20° for angles of attack from 
0° to 56°. These data were obtained at the same dynamic pressure and 
center-of -gravity location as for the longitudinal stability and control 
data. Presented for comparison with the free -flight-tunnel data are 
higher Eeynolds number data obtained from tests conducted at Convair. 

The Convair data are presented for a center-of -gravity position of 
27.5 percent of the mean aerodynamic chord. 

Basic design .- The lateral-stability characteristics determined from 
the free -flight-tunnel tests are presented in figure 6 for the basic 
configuration^, and a comparison of these (FFT) data and the Convair data 
is presented in figure 7* All free-fli^t-tunnel data are presented for 
an elevon deflection of -15°^ which corresponded approximately to the 
deflection needed to trim at hi^ lift coefficients. (See fig. 4.) The 
data of figure 6 show that the variation of the yawing-moment coeffi- 
cient and the rolling-moment coefficient C^ mth angle of side- 

slip p is fairly linear up to an angle of attack of 20° for the model 
■(■ath vertical tail (fig. 6(b)). At an angle of attack of 24° the tail- 
off configuration (fig. 6(a)) shows a large increase in directional 
instability. This increase in negative slope of the yai-n.ng-moment curve 
for the tail-off configuration is also reflected in the data for the 
tail-on configuration at an angle of attack of 24°. At angles of attack 
of 26° and hi^er, the tail-on data show a destabilizing break in the 
yavring-moment curve at sideslip angles greater than approximately ±5°. 


A con^arison of the data of figure 7 shows that^ at low and moderate 
angles of attack, the yawing-moment and rolling-moment curves for both 
models had the same general characteristics. At high angles of attack, 
however, the Convair data indicated less directional instability than 
the free-f light-tunnel data. 


The data of figures 6 and 7 summarized in figure 8 in terms of 
the side-force parameter directional-stability parameter 

and the effective-dihedral parameter • Since the data of figures 6 


and 7 nonlinear for some conditions, the data of figure 8 are pre- 
sented at low angles of sideslip (p = i2°) and hi^ angles of side- 
slip ( P = ±10°) . These data indicate that the free-fli^t-tunnel model 



MCA RM SL55B21 


7 


had lower directional stability over the angle -of -attack range than the 
Convair model and also became directionally unstable at a lower angle 
of attack than the Convair model. Because of the nonlinearities in the 
yawing-moment curves^ the directional stability determined for p = +10° 
decreased to zero at an angle of attack about 2° lower than that for 
3 = +2°. The effective dihedral positive for both models 

over the angle-of -attack range, ’(•rith the Convair model having higher 
values of hi^er angles of attack for p = +2°. 

!The yavmng-moment data in figure 8 for the free-flight-tunnel model 
are shovm for a center-of-gravity position of 27-5 percent c as well 
as of 50.0 percent c in order that a direct comparison may be made 
\rith the Convair data. The data indicate that changing the location of 
the center of gravity of the free-flight-tunnel model from 50.0 to 
27-5 percent of the mean aerodynamic chord had only a sli^t effect on 
the directional stability. Changing the center of gravity from 50.0 to 
27.5 percent of the mean aerodynamic chord had a negligible effect upon 
the rolling moment of the model. 


3 


The variation of the lateral-stability parameters Cy^, C^p^ and 
id.th lift coefficient and angle of attack for the F-102A are com- 


pared in figure 9 ^■rith data from reference 1 for the IF-102. In general 
the variation of the lateral-stability parameters ^-rith angle of attack 
was similar for the two models. Becaxise of the difference in lift 
curves for the P-102A and the YF-102 at high angles of attack (fig. 5)^ 
the plots of Cnp against lift coefficient (fig. 9) show that the 

directional stability drops off less abruptly for the F-102A than for 
the YF-102. The effective dihedral positive for both models 

over the lift- coefficient range >;lth the F-102A model having slightly 
higher values of -C^p at the higher lift coefficients. 


Modified design .- In an effort to obtain satisfactory static lateral- 
stability characteristics at high angles of attack, force tests were made 

of the model "vrith increased vertical- tail size ( — = O.I8) and id-th 

Vs / 

leading-edge slats. (See fig. 2.) The data obtained in these tests are 
presented in figures 10 and 11. The data of figure 12 compare 'the 
lateral-stability characteristics of the basic model with those of the 
modified model at angles of attack of 24° and 50°. The data of figures 10 
and 11 are summarized in figure I5 in terms of the lateral-stability 


parameters Cy^ 


:„p, and -C2p 


for angles of sideslip of +2° and +10°. 


The data of figure 12(a) show that at an angle of attack of 24°, 
increasing the size of the vertical tail = 0.l8^ caused the model to 


8 


MCA RM SL55B21 


"become directionally static and also made the curve linear. For the 
basic tail the leading-edge slats produced a small increment in direc- 
tional stability at low angles of sideslip and a very large increment 
at large angles of sideslip so that the overall result was a fairly 
linear variation of the ya^rLng-moment coefficient ^'rLth angles of side- 
slip. The slats had little effect on the directional stability when 
they were used in combination with the enlarged tail. 

The data of figure 12(b) show that;, at a = 50°^ model ■S'fith the 
basic tail was directionally unstable and the model with the enlarged 
tail was about neutrally stable for sideslip angles of ±5°- The model 
with either tail and without slats had a sharp destabilizing break at a 
sideslip angle of about +5°* ibe slats caused a destabilizing effect 
for small angles of sideslip and a large reduction in the directional 
instability for large angles of sideslip. The effects of increased tail 
size and leading-edge slats are shc^ more clearly in the summary data 
of figure 13 . The aileron and rudder control effectiveness for the 
basic model are presented in figure lit. 


FUGBT TESTS 


Flight tests vrere made from a lift coefficient of about O. 7 O throu^ 
the stall in order to determine the dynamic stability and control char- 
acteristics of the model in its basic configuration and \fith increased 
tail size and leading-edge slats. Flight tests were made at a center- 
of -gravity position of 27*5 percent c. Id^t ■(■ring loadings were used 
in order to minimize damage to the model in crackups. The model was 
floTO with coordinated aileron and rudder control and ivlth aileron-alone 
control. Aileron deflections of ±15° and a rudder deflection of ±25° 
were used for all conditions. Only relatively low-altitude conditions 
were simulated. 

The model behavior during flight was observed by a pilot sitiiated 
just aft of the tunnel test section. The pilot's observations and 
supplementary data obtained by motion-picture records served as a basis 
for all discussion of the flight tests. 


FUGHT-TEST RESULTS AND DISCUSSION 
Interpretation of Fli^t-Test Resiilts 


In interpreting the results of the model flight tests in terms of 
the full-scale airplane, it is necessary to consider any differences 
bet\reen the static stability derivatives of the model and those of the 


g W W !IiiIDliM3IA.5 



NACA RM SL55B21 


9 


full-scale airplane and any differences "between the scaled-up mass 
characteristics of the model and the mass characteristics of the air- 
plane. If there are no differences in these factors, then the airplane 
woixld he expected to ejdiihit dynamic characteristics similar to those 
of the free-flight-tunnel model. 

Althou^ no mass data were available for the full-scale airplane, 
the data of reference 1 show that the values of the scaled-up moments 
of inertia for the model of the Convair YF-102 were generally similar 
to those of the airplane at normal gross wei^t. Therefore, the mass 
data presented in table I for the scaled-up moments of inertia for the 
model of the F-102A are expected to be representative of those of the 
airplane at normal gross weight. It has been shown that the static 
stability characteristics of the free-flight-t unn el model at low 
Reynolds number are in fair agreement with the characteristics of the 
Convair model at higher Reynolds number. It is likely, however, that 
the abrupt changes noted in the stability parameters at high lift 
coefficients ■vri.ll occur at somewhat higher lift coefficients for the 
airplane than for the model. The dynamic behavior of the airplane is 
therefore expected to be similar to that of the free-flight-tunnel 
model, except that corresponding dynamic behavior might occur at higher 
lift coefficients. 

It should be pointed out that the full-scale airplane should be 
easier to fly than the model because its angular velocities will be 
only about one-third as high as those of the model. Another factor 
which should facilitate the pilot’s control of the airplane is the fact 
that he has independent aileron and rudder control rather than the 
coordinated aileron and rudder control which was used on the model. 

In interpreting the lateral-control characteristics of models in 
temns of full-scale airplanes, it has been found necessary in some cases 
to consider the differences in piloting technique between the models and 
the airplanes. A free-fli^t-tunnel study has revealed that airplanes 
which have high yavring inertia and low rolling inertia, such as the 
F-102A, tend to execute a pure rolling motion about the principal longi- 
tudinal axis of inertia, at least dirring the early stages of a rolling 
maneuver. VJhen these airplanes roll in this manner, an adverse sideslip 
angle about the stability axis is produced which is approximately equal 
to the angle of inclination of the principal axis times the sine of the 
angle of bank (t) sin . For instance, for a given angle of inclination 
of the principal axis of 20°, an airplane of this type when banked 50° 
vrill have an angle of adverse sideslip of 10° about the stability axis. 
Since the pilot of a free-flight-tunnel model flies the model from a 
remote position and can perform only very limited maneuvers, he does not 
object to the model’s executing essentially pure roll about the prin- 
cipal axis and apparently cannot detect the resulting adverse sideslip 
about the stability axis that might be objectionable to the pilot of the 





10 




MCA EM SL55B21 


»••• 
• • 

: : 

••• 

• • 
• • 


full-scale airplane. The estimation of the adverse sideslip character- 
istics of the airplane "based on the model flight tests is therefore 
expected to he optimistic. 

In the discussion of the fli^t tests, it should he pointed out 
that what the pilot observes is the yaw of the model in the tunnel, 
except in cases of violent motions of translation. This in effect is 
the same as sideslip, except for sign. In the discussion that follows, 
the terms yaw and sideslip are used interchangeably, yaw being used to 
imply that the attitude in the tunnel is the significant thing at the 
time and sideslip being used when the aerodynamic effects of sideslip 
are under consideration. Similar considerations apply to the usage of 
angle of attack and pitch. 

The results of the present investigation are illustrated more 
graphically by motion pictures of the flints of the model than is 
possible in a ■Hritten presentation. For this reason a motion-picture 
film supplement to this paper has been prepared and is available on 
loan from the MCA Headquarters, Washington, D. C. 


Longitudinal Stability and Control 

The longitudinal stability and control characteristics of the 
Convair F-102A were similar to those of the Convair YF-102 and were 
considered satisfactory for all conditions investigated. Although the 
longitudinal characteristics of the model were considered to be 
generally satisfactory, some diffic\ilty was encountered in flying the 
model in the high lift-coefficient range because of the large variation 
of drag •irLth lift, which is generally a characteristic of 1 ot 7 aspect- 
ratio delta tjlngs (ref. 5) • This large variation of drag with lift 
caused large variations of the glide angle "VTlth lift coefficient and 
necessitated almost continuous corrections to tunnel angle and airspeed 
in order to maintain fli^t in the txinnel. 


lateral Stability 

Basic design .- In general, the lateral stability characteristics 
for the basic configuration of the F-102A were similar to those for the 
basic configuration of the YF-102 tested in reference 1. At angles of 
attack below about 25° the model was easy to fly and the lateral stability 
was considered satisfactory. The lateral (Dutch Roll) oscillations were 
well dandled for all fli^t conditions. The directional stability decreased 
■^/Ith increasing angle of attack, and at an angle of attack near the stall 
(a = 26°) the model became directionally divergent. A typical fli^t 
record of the model at an angle of attack of 26° is shorn in figure 15(a-) • 
The model could be flown at this angle of attack as long as the pilot was 


^BOirra B EmaiL 



MCA EM SL55B21 


11 


aBle to keep the angle of sideslip small. It appeared, however, that, 
once an angle of sideslip of approximately 5° reached, the model 
could not he recovered and it diverged rapidly to larger angles of 
sideslip and snap-rolled into the tunnel wall. The directional diver- 
gence of the free -fU^t- tunnel model was evidently caused hy the large 
values of the static directional instahility at the hi^er angles of 
attack. The increased rate of the divergence at the moderate and large 
angles of sideslip is attributed to the sharp destabilizing break in the 
yawing-moment curve which occurred at the hi^er angles of attack. 
Another factor which might have contributed to the directional diver- 
gence was the decrease in positive effective dihedral in the hi^er 
angle -of -attack range. 

As flights were attempted at angles of attack above 26°, it became 
more difficult for the pilot to keep the model at small angles of side- 
slip and the divergence became more violent. Flights attempted at an 
angle of attack of 30° were very short because the model diverged soon 
after take-off. A flight record of the model at an angle of attack of 
approximately 30° is presented in figure 15(a) . For this case the model 
sideslipped to an angle greater than 30° and rolled to an angle of about 
30° before crashing into the tunnel wall. More effective xise of the 
rudder yawing moment co\ild probably have been obtained if the riiidder had 
been deflected independently, but even the maximum available yawing 
moment of the rudder would be insufficient to balance out the ya^fing 
moment due to sideslip at sideslip angles greater than approximately +5° 
at an angle of attack of 30°. 

In con^jaring the force-test data of figure 8, it is seen that the 
free-fli^t-tunnel model becomes directionally unstable at an angle of 
attack approximately 5° lower than that for the Convair model} this 
difference \ras probably partly caiised by differences in Reynolds number. 
Since the flight tests showed that the free-flight-tunnel model could 
be floim at an angle of attack about 5° higher than that at which 

became negative, it is possible that the airplane, because it will be 
operating at much higher Reynolds numbers, might not experience a 
directional divergence before it stalls. Other factors that might 
influence the high angle-of -attack behavior of the full-scale airplane 
are its slower yairing motions and independent rudder control which might 
enable the pilot to control the yawing motion fairly well and prevent a 
divergence in most cases, even at high angles of attack. The danger of 
a directional divergence >rill still exist, hcnrever, since the airplane 
might inadvertently reach the divergent conditions if the pilot becomes 
engrossed in some action, such as an evasive maneuver in combat. 


Modified design . - 
8o percent ~ 0-l8^ 


Increasing the size of the vertical tail by 
did not eliminate the directional divergence but 


did increase the angle of attack at which the divergence occurred. 



12 


MCA RM SL55B21 


Flights were obtained at angles of attack up to about 33° ''■rf-th the 
enlarged tail. Men flints were attempted at an angle of attack of 
35° hi^er, the model diverged in sideslip but the divergence ims 
less violent than with the basic tail at lower angles of attack. Records 
of the model with the enlarged tail are presented in figure 15(b) for 
model angles of attack of approximately 30° and 33°- 

The addition of leading-edge slats did not increase the angle of 
attack at which the model became directionally divergent. In fact, it 
appeared that the model diverged at a slightly lower angle of attack 
•^-Tith slats on than with slats off. The force -test data of figure 13 
show that, for sideslip angles of i2°, the addition of the slats increased 
the static directional instability at angles of attack greater than 
about 28°. A fli^t record showing a directional divergence of the 
model for an angle of attack of approximately 33° is presented in fig- 
ure 15 (c). The fact that the use of slats in combination with increased 
tail size failed to eliminate the directional divergence of the F-102A 
model in the high angle-of -attack range as they had done for the ^-102 
can be explained by the force-test data of figure 16. These data are 
for the t^ro models with the large vertical tail and leading-edge slats 
and show that, at the hi^er angles of attack, the F-102A had much 
greater static directional instability and much lower effective dihedral 
than the YF-102. This slat configuration, which had been selected on 
the basis of exploratory force tests on the YF-102 model, apparently 
\Tas not satisfactory for use on the F-102A model. A more suitable slat 
configuration for the F-102A could probably have been found from addi- 
tional exploratory force tests with this model, but such tests were 
considered beyond the scope of this investigation. 


lateral Control 

The lateral control characteristics of the basic and modified con- 
figurations were considered satisfactory over the lift- coefficient range 
investigated and were generally similar to those obtained for the YF-102 
model in reference 1. Althou^ the control characteristics could not be 
evaluated throu^ the stall for the basic configuration, it is believed 
that they would be similar to those of the model ’t-rith increased tail 
size. Ill flints of the model idLth increased tail size near the stall, 
some adverse sideslip with aileron alone was obtained because of the 
adverse yairing moments due to aileron deflection (fig. 14(a)). This 
adverse sideslipping was eliminated, however, by using the rudder in 
combination with the ailerons for coordinated control. In the higher 
angle-of -attack range, the model could be controlled satisfactorily 
until a directional divergence occurred. 

As previously pointed out, full-scale flight test of airplanes 
which have high yaid-ng inertia and low rolling inertia, similar to the 



mCA RM SL55B21 


15 


• ••• 
• • 

• • 
»• • 

• • 
»• • 

• ••• 


F-102A, indicated more severe adverse sideslip characteristics than were 
demonstrated hy the models of these airplanes in the free-fli^t tunnel. 
It is possible, therefore, that the adverse sideslipping behavior of the 
full-scale airplane may be objectionable at the hi^ angles of attack. 


CONCLUSIONS 


Eesxxlts have been presented from a free-fli^t-tunnel stability and 
control investigation of a l/lO-scale model simulating the Convair F-102A 
airplane. The model was flora throu^ a lift-coefficient range from 0.7 
to the stall, and only relative low-altitude conditions were simulated. 

From the results, the foUorang conclusions were drawn; 

1. In general, the fll^t characteristics for the basic configura- 
tion of the Convair F-102A airplane model were similar to those for the 
basic configuration of the Convair IF-lOE airplane model previously tested. 

2. The longitudinal stability characteristics were considered satis- 
factory for the basic and modified configurations over the lift-coefficient 
range investigated. 

5 . The lateral stability characteristics for the basic configuration 
were considered satisfactory over the lift-coefficient range investigated 
except near the stall where large values of static directional instability 
caused the model to be directionally divergent. 

it. An 80-percent increase in vertical- tail area Increased ihe angle 
of attack at which the model became directionally divergent. 

5 . The use of leading-edge slats in combination with an 80 -percent 
increase in vertical-tail area did not eliminate the directional diver- 
gence throu^ the stall on the F-102A model as they did for the 
YF-102 model. 

6. The longitudinal and lateral control characteristics were 
generally satisfactory. Although the adverse sideslip characteristics 
for the model were considered acceptable over the angle-of -attack 



• • 




JLU XO 


14 


mCA EM SL55B21 


range, analysis indicates that the adverse sideslip characteristics of 
, the full-scale airplane may he ohjectionahle at high angles of attack. 

• • 

• • 

■ • Langley Aeronautical Laboratory, 

'•** national Advisory Committee for Aeronautics, 

langley Field, Va., February 7# 1955* 



Peter C. Boisseau 
Aeronautical Research Scientist 


Thomas A. Harris 

Chief of Stabilliy Research Division 


JKS 


REFERENCES 


1. Johnson, Joseph L. , Jr., and Boisseau, Peter C. : Investigation of 

the Low-Speed Stability and Control Characteristics of a 
l/lO-Scale Model of the Convalr YF-102 Airplane in the Langley 
Free-Fll^t Tunnel. MCA RM SL55HA-, U. S. Air Force, 1955- 

2. Shortal, Joseph A., and Osterhout, Clayton J. ; Preliminary Stability 

and Control Tests in the MCA Free-Fli^t Hind Tunnel and Correla- 
tion With Full-Scale FU^t Tests. MCA TN 8l0, 1941. 

5* McKinney, Marion 0., Jr., and Drake, Hubert M. : Flight Character- 

IstiQS at Low Speed of Delta-Wing Models. MCA RM L7K07, 1948. 


MCA EM SL55B21 


15 


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TABJE I 

SCAIED-UP MASS AlH) DBOTSIONAL CHARACTERISTICS OP A 
1 / 10 -SCAIE MODEL SIMOLATHTG THE COWAIR F- 102 A 
AIRPIAlffi TESTED DT THE LMTGLEI 
PREE-PUGHT TUKHEL 


Weight, Ib 19 , 1*^0 

Wing loading, W/S, Ih/sq ft 27 - 5 ^ 

Relative density factor, 

Moments of inertia: 

l x, Slug-ft 2 17,580 

ly, slug-ft2 119,300 

I2, slug-ft^ 124,600 


Ratios of radii of gyration to wing spaa: 








0.1426 

0.572 

0.3801 


Wing: 

Airfoil ..... 

Area, sq ft 

Span, ft 

Aspect ratio 

Root chord, ft 

Tip chord, ft 

Ifean aerodynamic chord c, ft 

Longitudinal distance from leading-edge of root chord to 

leading edge of c, ft 

Svreephack of leading edge, deg 

Sweepforward of trailing edge, deg 

Dihedral, deg 

Incidence, deg 


IIACA 0004-65 (modified) 

695 

58-134 

2.09 

35.628 

0.801 

23.72 


11.96 

60 

5 

0 

0 


Slats; 

Span, percent wing span (two) 5I.7 

Chord, ft 1.36 


Elevens ; 

Area behind hinge line, percent wing area (two) 

SJpan, percent wing span (two) 

Chord, parallel to fusel^e reference axis: 

Root, ft 

Tip, ft 

Basic tail = 0 . 10 ^; 

Airfoil section 

Exposed area, sq ft * 

Span, ft 

Aspect ratio .. .......... 


Enlarged tail 


If 



Es^osed area, sq ft 

Span, ft 

Aspect ratio ... 


Rudder (same for both tails) : 

Area, sq ft 

Span, ft 

Root chord, ft 

Tip chord, ft 


10.12 

69-0 

3-15 

2.04 


UACA 0004-65 (modified) 

68.2 

8.66 

1.1 


118.5 

11.54 

1.12 


10.54 

5-72 

2.09 

1.59 



MCA EM SL55B21 





y 




Figiire 1,- The stability system of axes. Arrows indicate positive direc- 
tions of moments, forces, and angles. This system of axes is defined 
as an orthogonal system having the origin at the center of gravity and 
in which the Z-axis is in the plane of symmetry and peipjendiciilar to 
the relative -V'ri-ndj the X-axis is in the plane of symmetry and perpen- 
dicular to the Z-axis j and the Y-axis is perpendicular to the plane of 
symmetry. At a constant angle of attack, these axes are fixed in the 
airplane . 




NACA RM SL55B21 


CONFIDENTIAL 



CONFIDENTIAL 


L-85596 

Figure 5*- Photograph of l/lO-scale model simulating the Convair F-102A 
airplane tested in the Langley free-flight t\mnel. 
















D O 


MCA EM SL55B21 



Figure 7»- Comparison of lateral statilily cliaracteristlcs of model tested 
in the langley free -flight tunnel and model tested hy Convair» 


g HH 






ee 










-16 -8 0 8 16 24 

iS.deg 


•24 -16 -8 0 8 16 24 

^.deg 


Figure 10.- Lateral characteristics of the model tested in the Langley 
free-fli^t tunnel. 5g = ^ = 0.l8. 





16 24 


-24 -16 -8 


(b) ^ = 0.10, 

O 


Figure 11.- Continued. 




• •• 



Figure 11.- Concluded. 






4 8 12 16 20 24 28 32 36 

a, deg 


of the model tested in the Langley 
. 6 ^ = - 15 °. 






mCA EM SL55B21 


• ••• 


• • 
• ••• 
»••• 

» •• 

e • • 

»e» e 

• ••• 

» # • 
«• • 

»••• 





15 20 25 30 36 

a ,deg 




)S Sr 

o 0 +20 

□ 0 -20 

o 2+20 

^ +2 -20 

-5 +20 

c, +5 .20 

Q -10+20 

o +10-20 

0 -15 +20 

o +15 -20 


(a) Incremental force and moment (b) Force and moment coefficients 
coefficients due to aileron due to rudder deflection, 

deflection. 5a = ^15°- 


Figure l4. - Control effectiveness of the model tested in the Langley 
free -flight tunnel. 6e = -15° J ^ = 0.10.