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NACA
RESEARCH MEMORANDUM
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NATIONAL ADVISORY COMMITTEE
F O R AERONAUTICS
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A flight test w a s condmAed at the FUght Test Station of the Pilotless Aircraft Raaeerch Division at Wsllopa Island, Va ., t o determine t h e longltudfml control e n d stabilfty characteristics of a 0.7-scale model of the Fairchild Lerk p i l o t l e s e aircraf't,with :- the horizontal wing f l a p s deflected 15O. The data w e r e obtained by the use of 8 talemeter and a l a p by p*r, tmcklng. The. reeults c ahow a n increase of effectiveness of the longitudinal contml i n
producing, nor?ml,accelerat€ons UP t o 8 Mach mmibar- of 0.75 where . '
t,his effectivenees gradually decreased becomlng negative a t a Mach
number of 0 .@. .Previous t e s t s . trith w3ng $&pa undeflected showed . .
an increase i n effectivema8 to 8 Mach number of 0.93 where a eudden, loss. of control occmed. The model: m a - Qamically etaBle
throughout 'the'.speed ran&. The data c o a f i m d the drag increeee
at the c r i t i c a l .Mach m b e r .for large -&n&c38 of .attack R S . indfcated in hi@;h-speed"wind-tumml t € S t 8 * The I W A was requested by the Bureau of Aeromntics, ??my Department, t o mke flight t e s t s of the FafrcWld Lark pilotless- aircraft configwation t o evaluate the longitudinal s t a b i l i t y and control characteristics a t hfgh subsOn$c speeds i n order to predict the behavior of the full-scale atrcraft. In order to obtain thie information 0.5-8Cab models, e x t e r n a l l y geometrically eimilar to the Fairchild Lark, were constructed and. are to be flown a t the Flight Test Station of the Pilotless Aircraft Research Dfviaion at Wallops Island, Va. This paper covera reaulte of the flight of the second model. The results reported herein pertain to the 1ongitGdinal characterietics of the Fairchild Lark con- figuration wfth the horizontal wing f h p e deflected down 15*. The model me flown wlth a p m p m n e d flicker-type deflection of the rudder elevators. This flight test wae conducted July 24, 1946 t time from Launching, seconds M free-stream Mack number free-stream etatfc pressure P free-stream d,vnemfc pressure E free-stmam t o t a l preeaure n o m l - f o r c e coefficient CN chord.-force coefffcient CC W weight of model, poUnas S horizontal wlng area, 2.725 square feet a longitudinal acceleration, feet per second per second normal acceleration, feet per second per second
an
acceleration of gravity, 32.2 feet per second per second Q deflection of mdder el?:rstors, degreea (trailing edge down is positive) specific heat ratio; value taken, 1.4 deflection of horizontal-wing flape, degrees NACA RM No. W28a 3 c The simplified 0.5-~cale model used in thZe investigation m e
externally geomtrically 8Mlar t o the f i l l - s c a l e Lark (mQ-1) of
the P i l o t l e s e Plane Division of the Fairchild Engine and A f q b e
- Corporation. A description of the 0.5-scale Lark models ie given
in reference 1 .
Figure I preeents' 8 t&o-view dr8wing of t h e model whereae the generel specifimtions of both the model and Full-scale a i r c r e f t a r e given in table I. P h o t o p p h s of the mCel asld rocket motor with blast tube are & o m in f i w e 2.
For t h i s flight the rudder-elevators were deflected from
epproximatev 0 ' ' up t o -9 in a programed movement to give EL
fli.cker-t,vpe oper&tion. Thie control motion wa8 in operation before the mdel left the launcher and all during the flight.
l?igure 3, a tail-vfew photograph of the model, S ~ O F T B the deflected" ' control surfaces d the end of the blaat tube.
where p m e taken aB thepressure at 888 level at the time of the test. The mdel reached a n a l t i t u d e of only about 500 feet.
The velocity of sound f o r t h e test conditimm waa 1141 feet per second.
The normal-ecceleration factor and the nornusl-force coeffi- cient were baee3 on a linear variation with time of the wlng loading from the take-off condition to the burnout condition.
A tlns history of t h e f l i g h t 'of a 0.3-scale model Fairchild lark with the horizontal-wing f h p e d e f l e c t e d l5O ia presented in figure 5. This figure is a conversion of the t s m t e r record i n t o curves of rudder-elovator defloction, Msch number, norm1 accelemtion, and longitudinal acceleration again& time from launching. The total elapssd flight time i ~ t s 9.60 secondfa. The maximam speed somesponded t o a; Machnumber of 0.9 (rocket burnout) at a time of 3 . 7 0 second8 a f t e r launchfng The daahed part o f the Mach nuniber curve uas obtained by integration of the 1ongitllil-tnrs.P-acceleratim curve. ' Referring to figure 5 , it, m y be Been that the n o m 1 accel- eration, d t h the umal.short-period oecillatione, followed the deflec-Lion of the rudder-elevators until. 8 Mach.number of 0.75
v a s reached (t 2.6) . With ci?pi70ximtely e .To up elevator;the
normlacl=eleration changed from a poeitive value of about 2Og at
M = 0.75 t o a negative value of 5g a t M = 0 .s. As the speed
decreaeed from M = 0 .% (t 4.2), the longitudinal control 7 5 8 8 gmdusUy reeatablished. Control m a filly restored when
the wee3 decreased t o M = 0.75 ( t : 5.3) . The reeulta ehox
an increase of effectiveness of the longitudinal control fn producing nomalacceleratipne up t o M = 0.15 where t h f s effec- tiveness gradually decreaeed becoming negative' at M = 0.89.
Prev1ov.B teste, reference I, wlth wing fl&y not deflected, showed m incl-eane in cffectivenees up t o M = 0.93 where a sudden loss of control occurred.
The a b i l i t y of the rudder-elevators t o produce norm1 accelera- t i o n i e shown i n figure 6 a8 a plot of m m l - a c c e l e r a t i o n (%'i I ' 1 *\;g>,
factor - - againat Mach number. The accelerations med
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t o determine the variation of the norrml-accekmtian factor w e r e obtained by fairing out the traneient oscillations due t o the abrupt rudder-elevatordeflectims. The normal-accelerationfactor , .
m e computed -"for up .Tdder-elevator deflections since at 6 , = Oo .-the nodel was not mmhme'ring. Also shown on f Q u r e 6 is the curve of aorml-accelem'tioh factors f o r %he fli&t teet
with 6f = 00 (reference 1) . The difference between the powar-on
and power-off curve8 is probably due t o the rapidly changing values of the a b p e of the pitching-moment c u m at high lift coefficiente in t h i s Mach nmbar region &a shown in high-speed wind-tumd t e a t s . .
of a 0.25-scale model (reference 2) . The difference between
the 6f = 0 ' curve and tha Bp = 13O curve includes. the effects : 1 of different' &nter-of-pvit;I; b c a t i o n s (approximate* 2 percent.. .
moan aerodyrwnic chord) and tino dff'fersnt value8 of the slope of . . . .
the pitching-mmnt curves for hi& bed low lift coefficients. In order to*.obtain the ror?nal accelektfon produced.per degree of ' rudder-.&evator deflection for any desirkd wtng loading, dfvide the norml-ecce%emtionfactor by this a loading. Far example, at M = 0.77, the follohrlng ccrrip&riaom may be ,made: . .
~um-sCale a i r c r a f t .
Center of chord) Referring to figura 6 it m y be men that for FI = 0.83 and greater, the l i f t capabilities are m a t e r without the tTing f l a p s def lscted than with a deflection of 15O. Figure 7 preeente the variatian of normal-force coefficient wtth Mach rwiber for the power-on flight period, The nraxirmun value of Cx obtained is near the s t a l l e d region for % e h r k confi@naticn &8 determined f r o m reference 2 .
The model in fU&t exhibite6 Qm&mic s.t;abilit37 throughout the epeed range &e the short-period oscil'latione iz&€ced by the abrupt elevator movements were al.lray8 w e d 8 6 a h m by the normal- acceleration curve fn flgure 5 .
Figure 8 preeents curnee. of chord-forca and normal-force coefffcients for the power-off decelerating part of the flight.
The results of high-speed w3nfi-tunnel tests of a 0.25-scale model (reference 2) show an increase in drag coefffcient at M = 0.75 at high w l e e of attack w3t.h the wing f l a p a deflected 1T0. A t a cormspdnding Mach number &e ehown in figure 5 , t h e decreaee fn the longitudinal acceleration at t h e high normal acceleration indicates the drag increase. Referring t o figure 8 the higher value of Cc above M =I 0.75 as compared to thoee at lower values of M , also indfcatesthe drag increase.
. ! & a flight test to determine the langitudbnal stabili%,y and control ohamcteristics- with t h e horizontal wing flaps deflected down 15O f o r the.' Faikchild Iark pflotlese aircraft wag conducted a t the TL@t.Test Station of the P91otle88 A i r c r a f t ResearchDtvfeion: at Wallops Island, VEL. From the reeulte of this flight test, the following general conclusion6 are indicated: 1. Increase of effectiveness of the rudder-elemtore ae longitudfml control ~ I I producing normal accelemtions up t o M = 0.75 where t h f s effectiveness gradually decreaaed
becomtng .negatiw at M = 0 .@. Previous tests wfth wing flaps
undeflected 8howed.an increase in effoctivenees up t o M = 0.93 where a sudden.losa' of control occurred.
2. The model exhibited Qmmic stabildtg- throughout the speed ,-e.
3 . The decrease in 1one;ltudinal acceleration a t high n o m 1 accelerations confirms the drag increase for 6f = l5O i n the vicirdty of M =I 0.75 at large angles of attack as indicated i n high-aged w i n d - t m e l tests.
Langley Memorial Aeronautical Taboratory National Advlsory Committoe f o r Aeronautics .
I;angley ' Field, Va .
David c . stone
Aeronautical Engineer 1. Stone, David G . , and Mitcham, G r a Q L. : FlighCu-Test Evaluation of the Lor@tudlnal Stabiliw and Control Charecterietics of O.5-Scale Piodela of the Feirchild Lark Pilotlees-AfrcrsSt Configur&ion. Model with W i n g Ffaps Pirot Dsflected. TED No. M E A 2387. R” Mli No. -, Bur. Aero., 1946.
2. SmLth, Worman F. : Hi*-Speed Teste of I/&-Scale Model of Navy Special Miaaile Eark. NACA MR No. L S O I , Bw.
Aero., 1945.
Full-scale Itern aircraft Fuselage: Over-all length, in. 164 W x F m U m dfameter, in. 17 Winga : Aspect ratio 3 -49 3 -kg T o t a l S F , i n .
14 37
Chord (constant), in. 21.2 10.6
Angle of incidence, deg 0 0 0 0 Dihedral, deg 0 0 Sweep, rleg Airfoil eection: H o r f z o n h l wing N A A 16-209
V e l % i c a l w3ng mcIi l6-009
Wing area (per pafr including 2 725 fuselage), eq f-t Tail aurfaces : Spen, in.
Chord (constant) , in.
7 -7 Angle of incidence, deg 0 Dihedral, deg
4s
. Sweep, deg 0 Airfoil section MACA 16-008 Total projected area, sa_ ft 1.813 Propulsion: Tspe rocket Uquid Powdsr Approximato thrust, XI 600 Appro-te duruttTon, aec 220 3 -7 -rT&e-off, 18.58 c.g. location, percent chord 20 t&rmut, rg .81 'Take-off, 127.4 Weight, 1 %
urnout, 9 . 9 b
(Take-off, 46.7 Wing loeding, lb/eq ft l l 0
Burnout , 36.6
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~- - ~.--- -- -v~D___cr___\~F_Ur.5 . v-- -V - ------ -- -- • •••• • ••• •• ••• •••• CONFIDENTIAL NACA LIlfIlL 477.!!1S Figure 2 .- Ph otog ra phs of O. 5- sca le model Lark and rocket motor with bl ast tube .
CONFIDENTIAL WATI O NAL ADVISORY CQ YW lTTEE FOR A ERONAUTICS LANGLEY W EWORIAL A ERONAU TICAL LA BOR AT O RY - L ANGLEY FIELD, VA.
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• • • ••• •• • • •••• • •••• • • •••• • ••• •• ••• •••• CONFIDENTIAL ~ (:) F ig ure 2. - C oncluded.
o :::s (:) NATIONAL ADVISOR Y CO MM ITTBE YO R ARRONAUTICS LANGLEY MEM OR IAL AIRO R1UTICAL L AB ORATO RY - LANG L EY YI ELD , VA .
CONFIDENTIAL •• • • • .
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• • • • Fi g . 3 NACA RM No. L 6J 28a CO NFIDENTIAL I ~ .. : I I I Fi g ure 3. - Ta il - view phot ograph of a O. 5 -scale mode l Lark .
CO NFIDENTI AL NATIO NA L AD VIS O RY COM MITT EE FOR AER ON AUTICS LAWOLlY OK _ORI AL llROWiUTIC AL LABORA TORY - L4~OL 'T ~I'LD. V A.
• •• • CONFIDENTI AL NACA RM No. L6J28a Fig. 4 Figu re 4 . - Photo gr aph of a O. 5- sca le mode l Lark on launc he r .
N1T IO " 1L AD VI SO RY CO M~ITTB R FOR l KRORAUTICS L AN GLEY M EM ORI AL AB RO RA UT IC AL LAB O RAT O RY - L ANGLEY YJILD, VA .
CONF IDENTIAL
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T~me f r o m Iounchmg, t, sec figure 5.- Concluded.
. 2 .3 x 4 . 5 . 6 . 7 .8 .9 I. 0
Mach number, M NATIONAL ADVISORY COMMITTEE Foll " I T l ~
Figure 6.- Varlatmn of normal acceleratmn wtfh Mach number for fhght wlfb +e150
f i g u r e 7. - Vawafjon of normal-force coeffmenf wrfh Mach number for t h e
power-on portion of the fllghf. y/ifh +=/5.”
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