Document
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RESEARCH MEMORANDUM
Restriction/ Classification Cancelled -L <
NATIONAL ADVISORY COM .MITTEE
FOR AERONAUTICS
WASHINGTON P3LEtXESS-AIZtCRAFT CONFI(URATI0N By David G. Stone F l i a t t e s t s were conducted at the Flight W e t Station of the Pilotless Aircraft Resoarch Division a t Wall026 Island, V a . , t o determine the longitudinal control and etabrlity charecteristics of 0.5-scale lilodels of the Fairchild Lark pilotless aircraft with the tail in line with the wings a d with the horizontal wing flaps deflected 6 0 ' . The data were obtained by the w e of a telemstsr and by redax tracMng.
The W A w a ~ requesW!. by the Bureau of Aeronautics, Navy DeparZanent, to fliat testa of the Fairchild Lark pilotleas- aircraft configuration t o evaluate the longitudinal e t a b i l i t y and control characteristics a t high subsonic spoede in order to predict the behavior of the full-scaleaircra2.t. I n order t o obtain t h i s information 0.5-acde models, extornally ,gmmetzically 6 l m i l a ? ? . t O the
Fairchild Lark, were cona4xucted and flown a t the FliQht Test
Station of the PZlotlese A i r c r a f t Research Division a t Wallops IelasLd, Va. Tbe results reported herein pertaln t o the longitudinal characteristics of the f ol1mTin.g configurations: (1) model with the tail surfacee in line wtth the win- and wing flapB not deflected, Etnd (2) model of tb sta.nde3.d configuration (dihedral of tail surfaces 45’) with the wing f l a p s deflected 6 0 ° .
The full-scale Fairchild Lark is flown a t constant a n g l e of attack. The l i f t 5ncrement.s for maneuvering are gained by deflection of the horizontal w i n g flaps, and the longitudinal control surf ace8 a r e used only a8 trimmers. In these model toater thecontrolsurfaces produced angle o f ‘ attack, but t e s t e with various wing-flap deflection6 provided data for an evaluation of the ef2ectivemss of the trimming control function. Tho models were flown with a programmed flicker- type deflection of the longitudAna1 trimming control surfaces.
free-stmam Mach nwnber f’ree-sixeam static pressure, pounds per q u a r e foot free -8-m t o t a l preasure, murids p r s q w e foot normd-f orce coefficient chord-force coefficfent rata of, chanw of pitching-mament cwfficient with angle of attack, per de p e of u9t coefficient w i t h angle of attack,
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rate of chm- of n o - acceleration with elevator deflection, per d e w e period of oscillation, eeconds m w n t of inertia about Y-axdsr slug-fee t2 w e i g h t of model, pounds horizontal ving area, 2.725 square f e e t wing chord, 0.883 f a o t longitudinal acceleration, feet per second per second normal acceleration, feet per second ger second acceleration OT @avity, 32.2 fee% per second per second .
b f l e c t i o n of horizontal. wing flaps, degrees deflection of ruddor elevators o r elevators, degrees ( t r a i l i n g edge dmn is positive) sgecific heat ratio; value taken, 1 . 4 The sim;jlified 0.5-scale models used in this inveatig&tion were externally geometrically sirnib t o the full-scale Itark (TAQ-1) of t 3 e Pilotless ?lam Dfvision of the Fairchild Engine and A i r p l a m Corporation.Deacrigtione of the 0.5-scale L m k models are given in references 1 and. 2.
Figme 1 presents the general aJ;Tangen;snt of t h e model with the tail surfaces i n the erne plane as the w i n g s . A photomaph of this model with rocket motor end blaet tube is shown i n f i g m e 2.
The t a i l - i n - l i m b a t s were accoqlished by rotating the tail section, by fastening the vertical control surfaces at Oo d e f b c t i m , and by connecting the servosystem t o the horizontal control surfaces which were then elevators. For this flight the elevators ware
4 - mAcA #M No. LV17
deflected from approximately - 1 1 ' to loo in proCF;rammd mwmant t ; O
give a f licLmr - t p operation. This control-surface motion was in operation before the model left launcher & all dur- the flight. For t h i s flight the WFng flap8 were not def l e c b d .
The model was ground-launched without a booeter on a zero- length launcher s e t at an a n & of 45O from level. A photograph of tihe model on the launcher is shown in fig- 3 .
The @ n e r d specifications of the model a8 compared t o the full-scale aircraft ~JXI given in table I.
Standard Configuration Model Figure 2 : pressnt~l the general arranepsmnt of t h e model representing t h e standard configuration. A photograph of this &e1 wit11 rocket m o t o r and b h s t tube is shown in f f , - u r e 5. Thie f l i a t was made with
We horizontal wlng flaps dejTlected so, asd t h e rudder-elevator8
were deflected from approxiaa.tely -90 to 6 O i n a progranmd flicker- : . type operatian. A detail photograph of the horizontal w w flap deflected down 60° is shown fn figme 6.
The data from the fliats were obtained by the w e of a telems.i;er, C N Doppler radar, and photography. The four-channel telemeters
e v e continuous simala of the lon&tudinal acceleration, n o m 1
acceleration, -act pressure, and control-surface deflection. The impact-pressure r e c o r d from the blemeter was reduced to Mach number by the following oquation:
r 7 -1
L f where p was talosn ae the pressure at sea level a . t the tim 'of the c tests. Since +&e models reached an a l t f t u d e of anly about 500 feet during the hia-meed region, no hrge errors in M a r b introduced by t a k h g p constant. Tho velocitr of sound f o r the tail-in-line t e s t s was 1136 feet per second a;nd P or the standaud-coIlffgUration t e s t s was U42 feet per eecond.
The no-1-acceleration factor and. the nomal-force coefficient w8re based on a linear vaslation with tims of the virq loading from the t a h - o f f corditinn t o the burnout condition.
T i m -History Records
Tail-in-line model. - A t i m e history of the f l i g h t of a 0.5-ecale
madel Lark with the t a i l i n line with the win;;a and S, 5: Oo 5s presented in figure 9. The tot& elapeed fli&t time w a ~ 40.8 seconds.
Only the f i r s t 8 seconds of -the night are presented. since no change 5n the recorded fliat characteristics waa noted until the compressed air f o r the servosystem was exgended a few seconds later. The m a x l m m sped obtained corro~~ponds to a Mach nmbr of 0.87, occurs a t a time of 3 . s eeconde after Lamchiw, and coincides with the buK19ng out of the rocket motor. The dashed. Mach nmber curve was obtaineh 'by intewation of the longitudinal acceleration with the i n i t i a l ZoLnt a t t = 2.4 where the data frm the t o t a l head and radar check oxactly. After t = 3.8 the tote.l-hea8 channel failod
.t;O record p r o p r l y , and the recor- tine of the radar was expended
at t = 3.5.
Referring t o figure 9 , it may be seen that the normal acceleration, with the usual 6hOTt-j?eriOd oscillations, followed the deflection of the elevators throughout tho a p e d r - . 3ositive normal accelerations
of 7g a n & newtin acceleratione of 38 were obtained T o r elevator
deflectiane ~f a p p o m t e b -lo* & U. , respct3Vsly. NO reverad
of the normal accelemtion WES eqerienced' for the s p e d r b encountered. The low maximum velocity as compared with that ehown in reference I can be attributed to soor rocket thrust as indicated bg a2 w 76 as c q a r e d with a2 z 9g i n pmvious teats.
Figure 10 presents the variation of no--force coefZicienk with Mach nmber for the y m r - o n flight period. Fi&;Lure ll presents curves of chord-force and nomnal-force coefficients f o r the gomr-off decelerating p a r t of the f l i g h t . A t tines where CN = 0 the C may be said t o be equivalent t o drag coefficient; hence a t t = g.79 .
(M = 0.61) the b a g cosffici$nt is G .oBg, decreasing to 0.033 at
t = 7.56 (14 = 0.73).
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s t m w a c o n f i m t i o n model (6p = ab). - A t l m e hlstorg of the
flight of a 0.5-scale model Lark of standard configuration f l a p deflected 60° I s presented in figure 12. The total elapeed flight t i m e was 17.9 seconds. As determined from visual and photoE;raphic observation, the model b e ~ a elow r o l l near t = 1.8 indicating that the ri&t wing flap loosened resulting in unl.sawn def lectione, and near t = 7.0 the right wing flap brolce off caustng a severe roll. Further recordconversion beyond the t W 3 ths f l a p broke off was considered umzocen&my. The mRx-lmum s2oed obtaimd c m s p o m i a t o a Mach number of 0.91, occurs at t = 3.78, and coincideswith the burnLng out of the roclzt motor. The daehed Mach nwnber curve was obtained by intewation of the l o n g i t u d i n a l acceleration. For this flight the total-head channel and the Doppler ra,dar failed to record pro-mrly .
Referring t o figure 12, it may be seen that the normal etccelemtion, with the usual short-period oscillations, followed the &eflection of the rudder-elevators throu&ou-t; the speed range. Althea the rigbt wing flap had loosened, n o & accelsrgtions of 30s were obtained
f o r a rudderelevator deflection of -9 . Alao,czfter the f l a p
looerened, coneiderable waviness occurred in the longitudinal acceleration cur=.
F i w e 13 presents t h e variation of normal-i"orce c m f f i c i e n t with hhch number f o r the power-on flight geriod, Figme 14 presents curvee of chord-f orce and normal-force cwff icients for the power-off decelerating part of the fli&ht.
L o n g i t w 1 Stability Evaluatiom of the s t a t i c longitudinal &ability were obtaiwd 'by analysis of the short-period oscillation induced 3s the abrupt movement of' the elamtors as doscribed i n reference 3 . The following equation was used t a determine the r a t e of c - of pitch--mmnt coefficient wlth angle of attack: The m i a t i o m of ceater of gretvity and moment of Inertia are dcra included i n the computation of z.
The values of 2 obtained are for the nodel-flight center-
of -gravity locations which ?or t h e tail-in-line confiGurat1on vguried
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f r o m 18.86- to l8.gl-percent chord azd f o r t h e staJlde3.d configuration varied from 19.34- to 18.29-percent chord a8 the rocket motor Burned out.
The values of theperiod P determinscl from fidwes 9 and 12 are presented in f i , - u r e 15 +o s h o w the variation of the period of oscillation with Mach number. The scatter of "clhe test points on figure 15 indicates L h e amount o f ' error i n determining P. Considerable
scatter is shown for the sf X 6 0 ' case. T h i s may be &w t o
loosening of the flap.
Figure 15 presents the atatic longitudinal stability, as c o q u b d using the above equation, as a f m c t i o n of Mach nmber. These curv-es indicate that as M increaaes, the sta;bili* increases weatly f o r both confignrations. With the tail in lins with the win@, the s t a t i c stability i s lese at l o w M&ch nm3ers but increams faster and is =eater as M increases ae cmpared with the tail interdigitated
with the win-. For t h e case of sf w 6 0 ° at high values of CN
the s t a b i l i t y 608s n o t increase as fast w9kh incraaeing Mach nmber as at low values of C a .
dCL -t;o
Bjr Wriw the value of the s l o p of the l W t curve -
da be 0.08 (reference 3 ) , and also includ3ng t h e m i a t i o n of the center of gravity, the neutra.l points ware computed. f o r the- condttione.
The neutral points, of course, do n o t fnclude the pmbable changes
3 w i t h mch n m b r . m e variations of w neutral pointe for
da the tail-in-line node& (sf = O o ) end the stmdd-configura;tion model (8f 60°) wLth M are given in figme 17. A g a i n the increase i n s t a b i l i t y i s indicated by the lwge rearward movement of t h e neutral 2oint as M increases above 0.70.
O n the full-scale W k the tafl control surfaces are wed f o r triiiiing the aircraft o a v , where= the 1 W t Lncrmenta are gained by wing-flap deflections or all-movable ~ r i n g a . In these nodel tests the control surfaces produced c h m s e of angle of att;tack, but tests with various wing-flap deflections provided the data for an evaluation of the abil3tr of the control surfaces t o trim the c&&.ne at high lifts. The abilitx of the 1onGtudiW control Burfaces to produce noma1 accelerations is presented i n figure 18 as a plot of noMnal-acceleration factor against Mach nm3er.
The normal-acceleration factor was determined by the t o t a l changs i n f o r the t o t a l change i n 6,. Thi8 msthod of e; determining the normal-acceleration factor eliminates the need f o r de^^^ the 6e required f o r % = o as was required f o r
k z
computation of the normal.-acceleration factor in reference 1. The values of the norixd-acceleration factor neglect the effects of the difference be-hmen normal force and lift force, end the rate of chanm of flight path with tims. The maxinrum variation of center- of-gravity locations between models is app?oX&mtsly 2-prcent chord.
The dlfference due to gower-on i n t h e S, w 600 case is probably
cawed by thrust miaalinernsnt with center of p v i t y . In order
to obtain the normal accelerations produced per degree of elevator deflection f o r aqv desired wflng LORding, divide t h e normal-acceleration factor by thedesired wing loading. For example, at M = 0.75 and W = U O , the following comparisons ma;. be znaite: It is evldent that placing the tail i n line with the win38 results i n an appreciable loss in t h e a b i l i t y of the elevators to produce normal accelerations. Also, wing-flap deflections of 600
show a reduction in an - ger - 6 , as canpared to Q = 15O tzp t o
M X 0 8. It may be noted thet at M = 0.73 the longttudiml s b b i l i t g is aEroxisately the ~ a m e f o r t h e tail-in-line Etnd t a i l -
interdigitated tests; therefore,the change in 2 pep S, must be
due t o a reduction in the effectiveness of the trimming control.
Also, since the s t a b i l i t y is less for the S, 6 0 ° conf'lguration, this again indicates a reduction i n control effectfv~nesa f o r trimmin6. This reduction in controleffectiveness mag be attributed to w i n g d B e f f e c t s upon the t a i l .
Aleo sham In f i w e e 9 anit 12, the production of normal acceleration lags the application of control deflection. For the b s t of the tail-in-line rnode1,the lag in the produced normal acceler- ation is of the order 09 0.10 to 0.15 second after application of
m a RM NO. ~ 7 ~ 1 7 _____ 9
the control. For t h e m o d e l with I h p s deflected 60° the lag is a~proxinatelg 0.10 second. These lag t h e s na.y be conrpared to values of 0 t o 0.05 second shown i n references 1 and 2 f o r the s k n d a r d
nodel w i t h sf = 00 and 150. This l a g i n the effectiveness of t h e
elevators may be due again to wake interference frm the w i n g . The magnitude of this aerodynamic lag is such a s t o seriously complicate the internal s t a b i l i t y of an autopilot eervoeyetem.
The f l i & t tests t o deteraim the lon@tUdinal stability and control characteriskics with the tail i n line wit'a the wings ( 6 , = 0 0 ) Euld w i t h the horfzontel w i n g f l a y deflected down 600 for the Tairchild T m k pilotless a i r c r a f t mre conducted at the F l i g h t Test Station of the P i l o t l e s s Aircr82.t; Resewch Division at Wallops Island, Va. Ikon theresulte of the fliat b e t s , tihe following mneral conclusions indicated: Placing the t a i l i n line with t h e wiws results i n a comfdemble reduction fn the ef"fsctiveness of the longitudinal trbming control.
This configuration f a statically stable with laz@ increases in the longitudinal stability occurring above M ?z 0.7. m e model exhibSted dynanic stabilitr throughout the s p e d ran@. The eerodynamic lag of the t r - control encountered in the tail-in-line configuration would make angle-of-attack stabilization ver: difficult.
Deflecting the horizontal wing flag8 60° w i t h the tail inter- digitated with the w w a result8 in a reduction In the effectiveness 02 the tr-g control &B colilpared t o + = 1 5 ' up to M z 0 .&I.
Deflecting tbe f l a y s 6 0 ° produces a conaid.era.ble increase in static
~ O ~ g i t u & i n a l o 6 ~ b i l . i t y at h i & Mach mkers. S W U r l y , with flaps
deflected Go , the aerodynamic lag of .the t r - control w o u l d make
angle-of-attack stabilization d i f f i c u l t .
mvia
CJB ISACA RM No. L7F17 1. Stom, Eavid G., and Mitchmn, Gredy L. : Flight-Itest Evaluation of the Lonrptudinal Stability and Control Characterfatics of 0.5-Scale Mode1.a of the Fairchild Lark Pilotloss -Aircraft Configuration. Model with Wing Flaps Not Deflected - TED No.
NACA 2387. M C A MR Eo. ~ 6 ~ 2 2 , B u r . Aoro., 1946.
2. Stone, Davdd G.: F l i a t - T e a t ,Fvaluation of the Lon@tudiaal Stability &nd Control Characteristics of 0.g-Scale Models of the Fairchild Lezk Pilotless-AircraftConfiguration. Model with W m Flaps Deflected 1 5 ' - TED No. NACA 2385.
NACA RM No. L6J28a, Bur. Aero., 1946.
3. Stone, David G. : Fll&t-TestEvaluation of the LongLtUilinaJ.
S t a b i l i w and Control ~ h a r a c t e r f s t i c s of O.3-Scal.e Models of t h e Fairchild I m k ?ilotlem-Aircreft Conflwation. Static Longitudinal Stability o f ' Modela with, W i n & Flap B f l e c t i o w O f Oo a33d 1 5 ' - TED NO. WCA 2387. HACA REiL Ho. L6Ll7a, B u r . Aero., 1946.
Full-soale I t e m airoraft 164 & 8.5 3 -49 3.49 3 . 4 9 74 ’ 37 37 21.2 1 0 . 6 10.6 0 0 0 0 0 0 NACA 16-209 NACA 16-209 =A 1 6 . 6 - 0 0 9 UCA 16-009 1 0 . 9 2 -725 48 24 24 - 15.4 7.7 7 . 7 . o 0 0 45 , 0 0 0 ~ A C A 16-008 ~ A C A 16408 HACA 16-008 T o t a l projeoted. C O W projected 1.283 1.813 7 -25 P o v b r Liquid P e r 600 lo00 1200 220 3 -8 3 -9 f”0ff 18.86 LBUzaout 18.x “ o f f 125.4 Weight, lb 1060 Burnout 97.9 Take-aff 46.0 U O Burnout 35.9
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Momnt of iraertia abbut Y - a x i s , Take-off 8.9 slug” 221 (approx.1 . . ..
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-, NACA LMAL 49239 Figure 2. - Photograph of tail-in-line model with rocket motor and blast tube.
NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS CONFIDENTIAL LANGLEY MEMORIAL AERONAUTICAL LABORATORY - LANGLEY FIELD . VA .
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CONFIDENTIAL Fig. 3 NACA RM No. L7F17
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Figure 3.- Tall-m-line model on launcher.
• AT IO NA L AOVISOWf COW .. ITTEE .. A£RONA UTICS CONFIDENTIAL LANG LEY "E"OR IAL AERONAUT IC AL L ABORATORY - LANG L EY FI ELD. VA .
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-.;:J NACA LMAL 47795 Figure 5.- Photograph of standard configuration model with rocket motor and blast tube, ~ aq NATIONAL ADVISORY COMMITTEE '0" AERONAUTICS CONFIDENTIAL
LANOLtY MEMORIAL AERONAUTICAL LABORATORY - LANGLEY FI.I n VA ·
a1 CONFIDENTIAL NACA RI No . L7F17 Fi g . 6 Figure 6. - Phooo.gr apn of ho riz onta l wing flap deflected 60 on standard configuration mod el.
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NATIONAL ADVISORY CO MM ITTEE FOR AERONAUTICS CONFIDENTIAL LANGLEY W(M OR'AL AERONAUTICAL LABORATORY - LANGLEY FI (tO . VA .
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Figure 8.- Launching of model Lark; standard configuration; Of = 60°.
HATfOIIIAL A. OVl $ORY CO .... IT TEE FOR AERONAUTICS LAHGL£Y MEMOR I AL AERONA UTI CA L L ABORATOR Y - L ANG LE V Flr l n VA CONFIDENTIAL . . . . . . . . . . . - . . . . . . . . . . .
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