Document
Restriction/Classification Cancelled XATIONAL ADTTISOR'II COMMITTEE F O R fURONA.ePTICS' m-EARCH M E M O R A N D U M f o r the Air-Xe.;er i e l Com.nil, A r m : FDL-CBS 'TIE EFFECTIYT'XESS AT HIGH SPEEDS O F A 20-PE;RCENT-CHOBD PLAIF? TRAILING-mG3 FLAP ON TRE NACA 6 ~ 2 1 0 AIRFOIL SECTION By Louis S. Stivers, Jr .
SUMMARY An analysis has been made of the l i f t - c o n t r o l effectiveness of a 20-percent-chord plain trailing-edge f l a p on the NACA 65-210 a i r f o i l section from section lift-coefficient data obtained a t Mach numbers from 0.3 t o 0.875. I n addition, the effectiveness of the p l a i n f l a p as a l i f t - c o n t r o l device has been cornpaxed with the corresponding effectiveness of both a spoiler and a dive- recovery f l a g on the INCA 65-210 a i r f o i l section.
The analysis indica.tes t h a t the p l a i n trailing-edge f l a p employed on the 10-percent-thick a i r f o i l a t Mach numbers as high a.s 0.875 r e t a i n s a t l e a s t 56-yercent of its low--speed l f f t - c o n t r o l , effectiveness, end i s s u f f i c i e n t l y effective i n l a t e r a l conP;rol a,pplication, assuming a r i g i d wing, t o provide adequate airplane r o l l i n g characteristics.
The plain trailing-edge flap, as compared t o the spoiler and the dive-recovery flap, appears t o afford the most favorable characteristics as a device f o r controlling l i f t continuously throughout the range of Mach numbers from 0.3 t o 0.875.
A t Mach numbers above those f o r l i f t divergence of the wing, e i t h e r a plain f l a p or a dive-recovery f l a p may be used on a thin airplane wing t o provide auxiliary wing l i f t when the airplane is t o be controlled i n f l i g h t , other than i n dives, art these Mach numbers. The choice of a l i f t-conti-01 device f o r t h i s use, however, should include the consideration of other f a c t o r s such as the 2 COWID- WCA RM No. A7A3.7 increments of drag m d pitching moment accompanying the use of the device, and the structural and high-speed aerodynamic character- i s t i c s of the a.irplane which is t o employ the device.
Among many effects of compressibility which have been found i n f l i g h t and i n the wind tunnel is a large reduction i n the effectiveness of conventional airplane control surfaces at veloc- i t i e s considerably above the a i r f o i l c r i t i c a l speeds. In some instances the effectiveness has been shown t o reduce t o nemly zero a.t high speeds, thus definitely limiting the maximum speed of controlled f l i g h t . In order t o determine whether t h i s reduction i n effectiveness is influenced by the type of control surface employed, various l i f t - c o n t r o l devices on relatively thin a i r f o i l s h v e been investigated a t high speeds.
The l i f t - c o n t r o l effectiveness of spoilers and dive-recovery flags used on thin a.irf0il.s has been reportea in references 1 and 2.
The spoilers became. decreasingly effective with increa.sing projection a.t high Mach ambers, and eAhibited characteristics which were such The dive- as t o promote e r r a t i c l i f t control a t high speeds.
recovery f l a p s also showed generally unfavora3le characteristics f o r use, other than emergency, z s l i f t - c o n t r o l devices at hi@ speeds.
Wind-tunnel data, presented i n reference 3 f o r a. plain trailing-edge f l a p on a modified M C A k e r i e s a i r f o i l 19 percont thick indicated t h a t the effectiveness of a plain f l a p used f o r l a t e r a l control on a thick a.irf'oi1 rapidly decreases as the Mach nwnber i s increased above the a i r f o i l c r i t i c a l Mach number.
In order t o provide information on the l i f t - c o n t r o l effective- nsss of a plain trailing-edge f l a q on a representative thin EACA &series a i r f o i l , tlne present analysis wa.s undertaken. For comparative purposes, increments of section l i f t , drag, and pitching- moment coefficients f o r the plain f 1 a ~ together with the correspond- ing chasa.cteristics of the spoiler a d the dive-recovery f l a p axe The presented f o r the rmge of Mach numbers from 0.3 t o 0.875.
aaalysis pertaining t o the plain f l a p was made using data from reference 4. Z'he e f f e c t of the differences i n r i g i d i t y of the wind-tunnel models a d the various pasts of an airplme hats not been considered i n the present anclysis.
C 2 section l i f t coefficient increment o r decrement i n section l i f t coefficient Ac 2 increment i n section drag coefficient increment i n section moment coefficient &bout quarter- A c ~ /4 chord point M free-stream Mach number ao section angle of attack, &epees 6f f l a p dePlec ti on, degrees section laterax--control-eff ectiveness parmeter, ~ b s o l u t e ho/n6f va.1ue of the r a t i o of equiva,lent change i n section zngle of attack t o cha.nge i n f l a p deflection acgle st a constant sectlon l i f t coefficLent The present anslysis of f l a p effectiveness was mna.de using aerodynamic data. obtained i n tho Ames 1- by 3 5 f oot high-speed wind tunnel from t e s t s of the N X A 65-210 a i r f o i l equipped trith a.
2 G ~ e r c e n t - c h o r d p l a i n f l a p . mese data. were obtained f o r speeds ~ t h a. corre- ranging from 0.3 t o approxima,tely 0 -3 Mach number (w'
sponding range i n Reynolds numbers frcm ~?~pproximately 1 > : l o 6 t o
2 x 106) f o r a i r f o i l angles of a.tta,ck from -2O t o 8 O and f l a p deflections from -60 t o 60. More precisely, the flap def lecttons i n Oegrees were fotmd to be -6.3, -4.9, -2.6, 0, 1.9, 4.6, and 6.3.
The l i f t - c oef f i c i e n t data f o r a Mach number of zpproximately 0.9 were not obtained at a. s&f i c i e n t number of a i r f o i l angles of a t t a c k t o permit t h e i r use i n the gresent a.nalysis . For t h i s recson, only data. f o r Mzch numbers ass high a s 0.875 appe2.r i n the figures.
I n order t o indicate tho effectiveness of the p1a.h f l a p as a.
lift-producing device, increments of section l i f t coefficient f o r each angle of f l a g deflection have been determined. These increnents were obtained throughout the Mach number range at a.irf o i l angles of a.ttack corresporzding t o l i f t coefficients of 0, 0.2, 0.4, O.6,and 0.8 Faired curves showing these increments f o r at zero f l a p deflection.
constant Mach numbers a r e presented i n figure 1 a.s a. function of f l a p deflection, The same Pncrements f o r constant f l a p d c f l e c t i ~ n cross- plotted a t each a i r f o i l angle of attack given i n f i g a r e 1 m e presented i n figure 2 as a function of Ma~h n~m3er.
The effectiveness of a later&l.--control device is not indicated Some parameter cmpLetely by increments of l i f t coefficient alone.
must be used which considers the changes i n a i r f o i l lift-cmve s l o p with changes i n control surface deflection. The commonly used pan+ meter h r o / ~ ~ , def ine0 as the r a t i o of the change i n a i r f oil-sectior, angle of attack t o the change i n f l 8 p deflection necessmy t o msin- t a i n a. constant l i f t coeff icfent, has b e e l adopted f o r use i n the present analysis. The v m i a t i o n of t h i s parameter with Mach number for the plain f l a p of the prosent report is given i n figure 3 f o r several moderate l i f t coefficients. For comparison, the vasiatiozl of the lateral-control-effocti.crcness panmeter with Mach number f o r a 20--percent-choyd p l a i n f l a p on a l+percent-thick modified IV~LCA 6 w e r i e s a i r f o i l 9 s d s o shown i n f i g m e 3 . The curve f o r the La.tter a i r f o i l and f l a p w a s obtaiizod from figure 43 of ref erencc 3 .
For the present report, values of AUO/GS~ were taken as the absolute value of the average slopes of the curve of section angle of zttbck versus f l a p deflection over a range of f l a p deflections from -6' t o Go, f o r a constant section l i f t coefficient, A graph ( f i g . 4 ) has been prepared which i l l u s t r a t e s the respective va,riations with Mach number of increments i n section l i f t coefficient with f l a p deflection f o r the pla.in f l a p and f o r the dive-recovery flap, and of decremente i n section lift coeff i- From the high-peed investigs- c i e n t with projection f o r s spoiler.
t i o n (two-dimensional) of a spoiler locarted at several positions on the upper surfa.ce of the NACA 65-210 a i r f o i l section, it appeared t h a t the 50-percent-chord location was the mosl; s u i k b l e investigated.
'Decrements of l i f t coefficient f o r various spoiler projectfons a t
t h i s location are shown i n f i g w e 4 f o r an a . i r f o i l w g l e of a t t a c k
corresponding t o a, l i f t coefficient of 0.2 a t zero spoiler projection.
Similarly, the increments of l i f t coeff i c i e n t f o r several dive- recovery f l a p deflections a r e a190 shown i n f igurc 4 f o r a correspond- ing a i r f o i l angle of attack and f o r tho dive-recovery f l a p 1oca.ted at the 50-percent-chord position, The high-speed investigztion (two-- dimensiona~) of dive-recovery f l a p s indica.ted that, of three f l a p locations on the lower sxrface of the l?ACA 65-210 a.irfoi.1, the 50-percenkhord position w a s a l s o the most suitable locrztion.
The changes i n section and pitching-moment coefficients coyresponding t o the increients (or decrements) of l i f t coeffi-
c i a n t shown i n f i p e 4 axe presented i n f i g w e s 5 and 6, respec-
tively, f o r tho same three l i f t - c o n t r o l ,devices.
RACA RM No. A7A17 CONF IDEnTTm 5 The dotted portions of c e r t a i n curves a p ~ e s ~ i n g i n figures 1 cad 2 , and of the curve of figure 3 f o r the 19-percent-thick a i r f o i l % r e used t o indicate t h a t some uncertainty e x i s t s regasding tlzc v a l i d i t y of these data obtained i n the v i c i n i t y of the wind-tunnel choking &.ch nmaber (0.9 at zero angle of attack f o r the EICPL 65-210 a i r f o i l model, and approximately 0.74 at zero angle of a.ttaclc f o r the 19-percent-thick s i r f o i l model) .
DISCUSS I O U A desirable l i f t - c o n t r o l device f o r use on a i r c r a f t wings or t a i l surfaces is one which has uniform effectiveness throughout the range of Ma.cli numbers a.t which the device i s expected t o be employed. Furthermore, i f an airplane is t o maintain c ontrolle& f l i g h t a.t Mach numbers above those f o r l i f t divergence 02 the wing (which 2xe generaally lower than those f o r l i f t divergence of the t a . i l ) , it rmst be possible t o compensa.te f o r the l i f t deficiency
of the wing a.t thcse Mach numkers . These two particulars a r e
considered i n the succeeding di.scussim both i n regard t o the plain f1a.p o f the present ana.lysis and i n regard t o the cornpasisor?
the.t follows . m e two-dimensiona.1 da.ta. presented here i n can
indicate, i n general, the asrodynamic e f f e c t s on an a.iypla.ne wing or tail resulting from the use of one of the l i f t - c o n t r o l devices.
It should be remembered, however, t h a t several other factors which not considered i n t h i s ana.lysis, such a,s the differences i n 2.r~ the aerodynamic chara.c t e r i s t i c s of the t a , i l a.nd wing, the downwash a t the tail, avld the elevator hinge-mment chara.cteristics, may greatly a f f e c t th.e over-all longi tudinal-s t a b i l i t y adnd - c o n t r o l cha.rac t e r i s t i c s of an a.irpla.ne i n f l i g h t , especially a.t high speeds Effectiveness of the Plain Flap as a. Lift-Producing Device The increments of section l i f t coefficient shown i n figures 1 and 2, which indica.te the effectiveness o f the f l a p ass a. l i f t - producing device, show t h a t the effectivsness increases somewhat with increase i n Ma,ch number reaching a, maximum at a March number apparently depending on the magnitude of the f l a p deflection and the a i r f o i l angle of a.ttack. The Mach numbers f o r which the increments of l i f t coefficient ase greatest correspond approxi- ma.tely, i n most ca.ses, t o the a.irf o i l l i f t-divergence l.~acli numbers I n the ran2e of Mach numbers given i n figure 8 of reference 4.
from those a.t trl-iich the maximum increments occur t o 0.875 Mach The minimum number the effectiveness decreases i n varying degree.
effectiveness indicated, however, is never l e s s than 50 percent of Although the data of figures 1 and 2 indicate t h a t a t low speeds.
appreciable variations i n the effectiveness of the pla.in f l a p f o r Mach numbers between 0.3 and 0.875, it is believed t h a t these variations w i l l not too seriously l i m i t the application of t h i s control device on 10-percent-thick r i g i d a.irfoi1 i n the said Mach number range.
Figures 1 and 2 f u r t h e r indicate the plain f l a p t o be capable of providing substantial increments of l i f t coefficient f o r small f l a p deflections Mach numbers above those f o r a i r f o i l l i f t divergence. 'i'he pla.in flap, then, used e i t h e r on a t h i n r i g i d air- plane wing or $ a i l remains effective as a lift-producing device a t speeds area-ter than those corresponding t o t3e wing or tail l i f t divergence, respectively.
Eff ec ttveness of the Flain Flap f o r LateraL Control The lif+-control characteristics of a plain f l a p ant high speeds a r e of further significance from the standpoint of the 1a.teral The lateral-control effectiveness of the control of an airplane.
pla.in f l a p of the present report can be eva.luated from - h e data of figure 3 which show the vaziation with Mach number of the section For any given 1a.tera.l-c ontrol-ef f c c t ivene s s parame t e r ibo/n6f .
airplane the magnitude of the parameter p b / 2 ~ (helix angle genera.-ted by the wing t i p of an airplane i n r o l l ) is d i r e c t l y proportional t o the airfoil-section lateral-control parameter A study of the variations &/AQ (assuming a r i g i d airplaae wing).
of L!ao/A.sf with Mach number w i l l , accordingly, correspond t o a.
of an airplane employing the air- study of the variations of pb/2'~ f o i l and l i f t - c o n t r o l device. Furthermore, whatever decrease i n the values of p b / 2 ~ a i t h increase i n 14a.ch number can be allowed f o r &n airplane, consistent w i t h the maintenance of adequate l a t e r a l control, can a l s o be all0.t~ed f o r the airfoil-section pa.rameter &/A% .
The data of f lgure 3 f o r tine NA-CA 65--210 a i r f o i l wit11 a p l a h f f a p show an appreciable varia.tion i n lateral-control effectiveness over a range of modera.te l i f t coefficients a f t 'nigh Mach numbers.
The only marked decreases i n effectiveness, however, appeas 'GO The begin at Mach numbers near 0.83 f o r low l i f t coefficients.
l a r g e s t decrease i n offectivenoss, f o r Mach numbers up t o 0.875, is indicated f o r zero l i f t coefficient where the effectiveness has reduced t o a, value which i s approximately 50 percent of t h a t I n a Navy Department specification f o r the shown f o r low speeds, s t a b i l i t y a.nd control chasac t e r i s t i c s of airplanes (ref ercnce 5 ) , NACA RM No. A7A17 CONFIDEMTIAL 7 p b / 2 ~ f o r adequate no reduction i n the minimum allowable va.lue of l a t e r a l control is permitted f o r indicated a.irspeeds up t o 300 miles per hour, but a two-thirds reduction is permitted f o r an increase A t a s i n indicated airspeed f ~ o m 300 t o 500 miles pey hour.
a l t i t u d e of 10,000 f e e t (an a l t i t u d e specified i n reference 5 a.t which compliance with these lateral-control requirements ase t o be demonstrated bp the airplane i n f l i g h t ) indicated airspeeds of 300 and 500 miles per hour correspond, respectively, t o approxima.te1y 0.7 and 0.8 Mach nmbers . The plain trailing-edge f l a p a.pplied t o z.
r i g i d wing appears, then, t o exhibit adequate lateral-control character- i s t i c s up t o Mach nmbers as high as 0.875.
A comparison of the curves of figure 3 f o r the two a i r f o i l s employing 20-percenbchord f l a p s shows tha.t the effectiveness exhibited by the f l a p on the 19-pe~cent -thick a i r f o i l at high speeds is quite different fronl tha.t f o r the 10-percent-thick a.irfoi1.
The curve f o r the 19-percent-thick a i r f o i l shows a marked decrease i n the effectiveness of the f l a p a.t a Mach number near 0.70 which is approxima.tely 0.13 Mach number l e s s than t h a t corresponding t o the abrupt decrease i n effectiveness of the f l a p on the 10-percen* thlck a i r f o i l a.t low l i f t coefficients. It can a.lso be noted from the dcta of figure 3 that, while serious losses i n the effectiveness of a, f l a p on a. 19-percent--thick a.irfoil can be expected above Mach numbers of thz order of 0.7, no severe losses should be expected for a p l a i n f l a p on a, 10-percent-thick a i r f o i l , especia.lly f o r higher l i f t coefficients, up t o Mach numbers approa.ching 0.875.
Comparison of the L i f W o n t r o l Effectiveness of a Spoiler, n, Divo-Rocovery Flap, m d a P1a.h Flap The rela.tive merits of a spoiler, a, dive-recovery f lap, and a.
a.irf o i l can be evaluated p l a i n fla,p f o r providing 1 i f t control on It can be from the l i f t - c o e f f i c i e n t da.ta presented i n figure 4.
seen rea,dily from the data t h a t the variations with Mach number of the l i f t - c on-L~ol effectiveness of the spoiler and the dive-recovery f 1a.p from a Mach nwnber of 0.3 t o 0.875 are considerably larger than the corresponding variations f o r the pla.in f l a p . Becmsc of these large veriations i n effectiveness f o r the dive-recovery flap, m d especially f o r the spoiler, an ai_:qlrzne control systern employ- ing e i t h e r of these devices would tend t o provide a t high speeds too rapid airplane response t o control movements i f setisfa.ctory low- For producing l i f t speed control chaxac t e r i s t i c s were maintained.
numbers, the plain continuously throughout a, wide range of Mach tra.iling-edge f la?, a.c c ordingly, appeays t o possess the most favor- ciblo char;2..cteristics, 8 C OlVFIDENTLAL NACA R M Bo. A7A17 For providing auxiliavy l i f t a t Mach numbers above those f o r a.irfoi.1 l i f t divergence, the p l a i n f l a p deflected i n a positive sense and the dive-recovery f l a p are considered f o r positive increments of l i f t ; whereas the plain f l a p deflected i n a negative sense and the spoiler, on the other hand, axe constdered f o r negative incre- ments of l i f t . The data of figure 4 show t h a t each of these l i f t devices is capable of providing increments (or decrements) of l i f t coefficient i n the range of Mach nwnbers between 0.75 and 0.875.
(This range includes Mach numbers a,bove those f o r l i f t divergence of Tliese increments, however, vary d i f f e r e n t l y f o r each the a i r f o i l ) .
l i f t device with changes i n Mach number and decrease with increase i n Mach nmber a t the highest Mach numbers shown, except f o r the 10' deflection of the dive-recovery f l a p and f o r positive deflections of the p l a i n f l a p . The plain f l a g appears t o ha.ve no p a r t i c u l a r advantage over the dive-recovery f l a p f o r providing positive incre- ments of l i f t a.t Mach ambers between 0.75 and 0.875 on a 10-percent- A t the Mach thick a i r f o i l unless it be a t the highest Mach nwnbers.
numbers near 0.875 the data f o r the plain f l a p show t h a t the increments of l i f t coefficient f o r the larger f l a g deflections do not continue t o decrease w i t h increase i n Mach number as the corresponding incre- ments do f o r the dive-recovery f l a p .
The increments of dra.g coefficient corrcsponding t o constant increments of l i f t coefficient, as shown i n figure 5 , axe seen t o be quite d i f f e r e n t f o r the three l i f t - c o n t r o l devices. The chara.cteristics f o r the plain f l a p appear t o be the most desirable, since the data indicate t h a t the increments i n &rag accompanying a given increment i n l i f t is the l e a s t f o r the p l a i n f l a p a t any Mach nwnber from 0.3 t o 0.875. Between 0.75 and 0.875 Mach numbers the increments i n drag coefficient f o r constant increments of l i f t coefficient of the dive- recovery f l a p increase very rapidly with increase i n Mach nmber.
I n the case where a 1if"ccontxol device is used on an airplane wing as a purely emergency implement f o r a i d i n recovery from high-speed dives, a substantial increase i~ drag, such as noted f o r the dive- recovery f l a p , may be desirable i n order t o l i m i t the diving speed of the airplane.
A t constant increments of l i f t coefficient, the increments of pitching-moment coefficient presented i n figure 6 do not w r y a @eat deal. with change i n k c h number except, f o r the lllost part, a.t the highest Mach nwnbers. For nega.tive increments of lift a t Mach numbers between 0.3 and 0.875, the pla.in f l a p and the spoiler exhibit, i n general, positive increments of pitching moment wbich tend t o increase a t the highest Mach numbers f o r the larger negative i n c r e ments of l i f t . The pitching-moment increments f o r positive incre- ments of l i f t a r e negative f o r the plain f l a p , and positive f o r the EACA RM No. A7A17 COB@' IDEI\YT IAL 9 dive-recovery f l a p except f o r the larger increments of l i f t a t high Mach m b e r s . The data show t h a t the increments of pitching moment a r e always more positive f o r the dive-recovery f1a.p than a r e the corresponding pitching-moment increments f o r the plain f l a p . I n the range of &,ch num3ers from 0.75 t o 0.875 %he pitciiing-moment coefficient^ f o r the plain f l a p are always, negative (not i n the direction t o oppose the diving tendency); whereas f o r t2e dive- recovery f l a p they appear t o be e i t h e r positive or negative, depend- ing on the Mach number and the increment of l i f t coefficient. A negative increment of pitching moment accompanying the use of any I i f D c o n t r o l device a t high subsonic speeds should certainly be consid- ered i n the structura.1 and aerodynmic design of an airplane tail.
Tlie diving tendency of a.irplanes, resulting from the l o s s i n wing l i f t a.t Mach numbers above those f o r l i f t divergence of the wing, is g e r i e ~ d l y accompanied by an increase i n longitudinal s t a b i l i t y a d by trim clzanges. A s a consequence, the co3*ol forces of some airpla.neisin high-speed dives increase t o such an extent t h a t it has been found necessary t o eraploy dive-recovery f l a p s a s an emergency device t o a i d %he p i l o t i n pulling out from the dives.
O n the other hand, p i l o t s of some of the more recent high-speed a i r c r z f t have effected recovery from high-speed dives without recourse t o emergency devices. In mlergcncy applications the dive-recovery f l a p s w e a.dvantageous i n t h a t tliey increase the wing l i f t f o r aLr- plane trim by providing an increment of l i f t together with a favor- Since the data of Tigure 6 f o r t ' n c dive- able pull-out moment.
recovery f l a p show t h a t the pitcbing-aonent increment is not jlrmys positive, it would appear that the us@ of these f l a p s on an airplane wing na$ not always provide favorable pitching moments f o r dive
recovery as Llie Mach number or increment of l i f t is increaood. he
dsta, of reference 2 show t h a t tke dive-recovery f i a p loca.ted on the a i r f o i l as far forward as the 30-percenkhord position also provides negative increments of pitching rnoment at high subsonic Mach nmhers, except when tlie f l a p has a. small chord r a t i o . ) If a n airplane is t o be control-led i n f l i g h t , other than i n dives, a t Mach nm]hers above those f o r l i f t divergence of the wing, the use of dive-recovery f l a p s a t these Na.ch nunbers t o provide e ~ u r i l i w y l i f t on the wing may be limited by the large increase i n dra.g. The choice of a l i f t-control devic.e f o r such operation should a l s o depend upon z, consideration of other f a c t o r s such as the incre- ments of pitching rnoment accompanying thc: use of tho devi-cc, 2nd the s t r u c t u r a l and high-speed aerodynamic chaxacteristics of the airplane which is t o eaploy the device, The analysis of the l i f t - c o n t r o l characteristics of a 2Qpercent- chord plain trailing-edge f l a p on -tho RACA 65-210 a i r f o i l section and a comparison of the effectiveness of t h i s device with W-at of both -&e spoiler and the dive-recovery f l a p indicate the follo~iing: A t Mach numbers as high as 0.877, the plain f l a p on the 1, 10-percent-thick a i r f o i l r e t a i n s tzt l e a s t 50 percent of its low- speed l i f t-control effectiveness, and is s u f f i c i e n t l y effective i n l a t e r a l contro1,assuming a r i g i d wing, t o provide adequate air- plane r o l l i n g characteristics.
A s compared t o the spoiler and the dive-recovery flap, the 2.
pla.in tra.iling-edge f l a p would appear t o afford the most favorable c h w a c t e r i e t i c s as a device f o r control3ing l i f t continuously throughout a range of Mach numbers from 0.3 t o 0.875.
An airplane employing t h i n wings which is t o be controlled 3 .
i n f l i g h t , other than i n dives, at Mach numbers above those f o r l i f t divergence of the wing ma.y use e i t h e r a plain f l a p or a dive-recovery f l a p at these Mach numbers t o provide auxiliary l i f t on the wing. It should be remembered, however, tha.t the choice of a devlce f o r this use should include the consideration
of other f a c t o r s such as the increments of d.rw and pitching
moment accompanying the use of the device, and the structura;l and high-speed aerodynamic characteristics of the airplane f o r wbich the choice is t o be made.
Ames Aeronauticd Laboratory, National Advisory Committee f o r Aeronautics, Moffett ~ i e l d , Calif.
NACA R M Wo. A7A17 CONE'DEPFTW 1 1 1, Velasco, Carlos E.: High-Gpeed Winlt-ifumel Investigation of Spoilers f o r Later& Control on the NACA 651-210 A i r f o i l i Section. NACA CMR NO. A'jKO2, 1945.
2. Olson, Robert N., a d Benoapa, Joseph N. : High-Speed Wind-- Tunnel Investigation of Dive-Recovery Flaps f o r L i f t Control i .
on the NACA 6 5 ~ ~ 2 1 0 A i r f o i l . NACA ARR 6~23, 1946.
Lindsey, W.F. : Effect of Compressibility on the P r e s ~ u r e s and 3.
65,3419 Airf oikHaving a Forces Acting on a Modified MCA
, L
&' d 0,2Whord Flap. 8ACA ACR No. L%3la, 1946.
C, 4. Graham, Donald J., and Adams, Charles N.: Wind-Tunnel Investiga- t i o n of a 20-Porcent-Chord Plain Flap on the NACA 651-210 i NACA CMR No. A T 0 5 , A i r f c i l f o r L i f t Control a n t High Speeds.
1945.
Specif'ica.tionf o r S-tability and Control Characteristics on. : Spec. No. LlPA, B u r . Aero., Navy Dept., of Airplanes, Apr . 17, 1945.
EACA RM No. A7M7 FIGURE m E M D S
Figure 1 .- Variation of the increment of section l i f t coefficient
w i t h f l a p deflection at vwious Mach numbers f o r several m g l e s of attack of the NACA 65-210 a i r f o i l with a 0.20-chord f l a p .
Figure 1.- Concluded. NACA 6 w 1 0 a i r f o i l with a 0 . 2 h h o r d plain f l a p .
Figure 2.- Variation of the increment of section l i f t coefficient with Mach nmber f o r various f l a p deflections and angles of attack of the MACA 65-210 a i r f o i l with a. 0.2O-chord p l a i n f l a p .
Figure 2 .- Concluded. NACA 65-2~0 a i r f o i l with a 0.20-chord plain
flap.
Figure 3.- Comparison of the l a t e r a l - c o n t r o l effectiveness at various Mach numbers f o r the NACA 65-210 m d l p p e r c e n t thick 6 ~ e r i e s a i r f o i l s with 20-percent-chord plain f l a g s , F i g w e 4.- Comparison of the l i f t - c o n t r o l characteristics of a.
spoiler, a dive-recovery flap, and a p l a i n f l a p on the NACA 65210 a m o i l section apt an angle of a.ttack corresponding t o a l i f t coefficient of 0.2 f o r zero deflection of the control device.
Figure 5.- Compari~on of the increments of section drag coefficient corresponding t o constant values of increment i n l i f t coefficient given by a spoiler, a dive-recovery flap, and a plaSn f l a p on the NACA 65-210 a i r f o i l section a t a- angle of a t t a c k corresponding t o a l i f t coefffcient of 0.2 for zero deflection of the ~ o n t r o l device.
Figure 6 .- Comparison of the increments of section moment coeff i--
c i e n t corresponding t o constant values of increment i n l i f t coefficient given by a spoiler, a. dive-recovery flap, and a. pfa.in f l a p on the NACA 65-210 a i r f o i l section a t a 1 1 angle of attack corresponding t o a l i f t coefficient of 0,2 f o r zero deflection of the control device.
Restriction/Classification Cancelled