Document
r
RESEARCH .MEMORANDUM
THE EFFECTIVENESS AT HIGH SPEEDS OF A 10-PERCENT-CHORD PLAIN TRAILING-EDGE FLAP ON T H E NACA 65-210 AIRFOIL SECTION By Richard J. Ilk A m e s Aeronautical Laboratory Moffett Field, Calif.
CLASSiFICATfON CAWCELL€
NATIONAL ADVISORY COMMITTEE
FOR AERONAUTICS
WASHINGTON June 14, 1948 TIIE: EFFECTIVENESS AT HIGH SPEEDS OF A IO-PEBCENT-CHORD PLAIN TRAILING-EEGE FLAP ON ' T H E NACA 65-210 AIRFOIL SECTION By Riahard J . I U S SuMEdaRY A high-speed wind-tunnel investigation h8s been made to determine the effectiveness of a 10-peroent-chord plain f l a p on the YACA 65-210 a i r f o i l sectton. For f l a p defleotionsranging f'rm approximately - 1 2 ' to 12O, section drag and l i f t forces were measured o'ver a range of Xach numbers from 0 . 3 t o 8bout 0.875 f o r angles of attack from -2O t o 8'0 Incraents i n e e c t i o n l i f t coeffioient are presented aa 822 indi- catLon o f the lift-producingcharacteristics of the plainflap. Kore signifioantly, values of the section flap-effeotiveness parsmeter are given as a measure of the effectiveness of the flap.
The t e s t results indicate that at no apeed within the Fnvestigated range does t h e l i f t increment f o r a given deflection of the 10-peroent- chord f l a p f a l l below 5 0 peroent o f its low-speed value, A aomparison o f the effectivenesa of the IO-percent-ohord f l a p with that of a 20- psrcent-chord f l a p indicateta that a reduction in flap-chord ratio from 0.20 to 0 . 1 0 lessens the severity of the effeotivmesa losaes at super- c r i t i c a l speeds. The 20-percent-ohord flap, hornever, remains more effective throughout the entire 'ddach nunber range of the present investi- gation.
Numerous wind-tunnel and f l i g h t t e s t s have indioated that conwm- tional airplane control surfaces experience a considerable loss i n ePfectiveness at high subsonic speeds. Although analytical methods appear reasonably accurate in predicting the low-speed effeotiveness of flap-type controls, reliable estimates of airplane oontrol oharacteristics at high Mach n m b e r s can only be made fra pertinent experhental data.
k t present, the availableexperimental data a r e Fnsufficient to allow quantitative estimates o f the high-speed variation-in control effectiveness with Machnumber to be made for arbitrary airfoil control-surface canbi- nations.
UNCLASSIFIED
2 NACA W N O . - 1 The prose& i w e e t i g a t i a a -8 undertaken t o provide Infonutian rn the oontrol effeotiveness of I p l a i n trailing-edge oantrol aurfaoe on a mpresentativm thin U C A d s s r i e r a i r f o i l . Roa a oamprrlsan of the variation in f l a p offeotirsnssu with &oh number for the 10-peroent-ohord flap of the pres& report- with riaoihr data previ- ously obtained fn the Am08 1- by S g f o o t high-speed r i n d tuptul for a ZO=perosnt-ohord flap on the U C A 6S-210 a i r f o i l , I t warn hoped that mame oonolusions regardkg the effeot of flap-ohord ratio on high- s p e d oontrol dmraoterirtior aould be made. The prseent rrulysia has not omsidered the effeotu of differenoes i n rig5dity whlohwould axirt on the l i f t i n g surfaoer of an aotual airplane.
SPMBOIB a i r f o i l section drag ooefficisnrt "d M , free-stream Maoh number airfoil seation angle of attaak, degrees 6f f l a p defleotion, degrees . A a & 8 f section flap-effectivsness parameter, absolutevalue o f the r a t i o of equivalent ohange in aection angle of attaok to ohango in flap-defleotion angle at a oonstant motion l i f t ooeffioiallt All tests were oonduotsd in the Ames 1- by *foot high-speed wind tunnel which is 8. low-turbulence, "dimensional-flow, cloesd- throat tunnel.
Seven models of 6-inoh ohord were oonstructed of s o l i d aluninum alloy to represent various deflections of a 10-percent-chord plain trailing-edge flap of true a i r f o i l oontour employed on the NACA 65-210 airfoil seotion. The actual flapdeflections in degrees were - 1 1 . 8 , 4 . 8 , 0 , 1 . 7 , 4.0, 6.9, and 10.6. The ordinates for the NACA 65-210 airfoil are given i n t a b l e I and a sketch of a typical profile is shown i n figure 1.
RESULTS AM) DISCUSSION !&e seotion drag and lift ohara&erietios of the XACA 65410 airfoil Kith a 10-percent-ohord p l a i n flap at mious defleotion angles a r e preeentad as I % funotion of I4d.1 nuniber in f3gures 2 and 3 , respeotively. The variation of imremgnt in s e o t i m lfft w e f f i o i e n t w i t h flap defleotfon is s h c n r r n in figure 4 for mrious W h n~.anbers and o m s t a n t angles of atfmk. L i f t laoremsntrs for oon8tm-b flap defleotfons hum been oross-pIott;ed for t h e same airfoil angles o f attaok g i n in figure 4 and are presented i n figure 5 as E A funotion of Maoh number.
The data of both figures 4 and 6 indicate that the liFt inorementa o f the IO-peroent+hord flap inorease .Rfith &wh ntmiber, reaohing a 4 MCA RM Nor - 1 maximLMl at a &oh nunfber dependent upon a f r f o i l angle of attaok and magnitude of the flap-defleotion angle. A t any bfach rider in the range f r a m 0 . 3 t o 0 . 8 7 S s the inorement in lift aoeffioienf produoed by the 10-percmt-ohord f l a p is nemer less than M) p e r o a t of i t s low-speed value for a gfven defleation.
For a p l a h trailfng-gdge flap, the oontrol effeotivensrs be evaluated f’rarrr data whidh clanonstrate the variation with lrfaoh number of theflap-effeotiveness parameter. This flap-effeotiveness parameter Aa&f is equal t o t h e absolutevalue of t h e ohetnge in seoticm angle of aftaok per unit ohange in f l a p defleotion at a oonertmt lifi uoeffioient. For the presentreport, uumss of seoticm angle of attack as a Function of f l a p defleution at oonstant lif’t coeffiuientwere plotted for v a r i o u s & a h nmnbers. The absolute value of the average slope of eaoh m, f r o m 6 f a-100 t o 6s 100, m s talcem as the f l a p e f f s o t i m e s s for 8 given lift ooeffioient and &oh number. The effeoti~eness paramster varies s l i g h t l y w-ith f l a p deflection and usually deweases as the deflestian angle iPare8ses.
The effeativmerrs of the 10-percent-chord plain flap operating at moderate lift aoefficients r a r i e s appreoiably over a rmge of Maoh numbers frcun 0 . 3 t o 0.876. (See fig. 6 0 ) A t the l m e t speeds the values of Aa a / A & for the lO-peraent-ahord flap are a p p r o d m t e l y 85 peroept of the theorstioal d u e (referenoe 2) for thin airfoils. The nap effeotivenese Beoresees graduklly with an inarease in &oh m b e r from 0.3 t o approxhmtely 0.776, after lihiah a more marked d m ~ e a 8 8 is exhibited. The large& reduction In f l a p effeati.crenese, over the M a & nurnber range frm 0.3 t o 0 0 8 7 6 ~ i13 hdioated from zero lift ooeffioientwhere the effectiveness kas deoreased t o about 65 peroent of its law-speed Value. The variation betmen the experimental and thsoretioal
values of Aa dA& be a t t r i b u t e dt o the lnfluenoe o f viseositys
the effects of mhioh are not oonsidered fn - & e t 2 1 e o r y . S h o e the
rate o f inorease in boundary-layer’ thiokiess w i t h flap defleotion is usually greater than the rate of inorease in boundary-layer thiokness w i t h angle of attaok (referenoe 2), the slope bot/66f is deoreased more by visaoeity than i a bq$arb.
Fram a amparison of the variation in flap eff’eoti-v-emse with Idach number for a lO-peroent-hrd and a 20-percent-uhord plain flap on the NACA 66-210 a i r f o f l eeotfon (flg. 7 ) * it oan be Been that the loss in flap effewtiveness at the highest Haoh number is oonsiderably less aevere for the smaller-ohordflap.Despite the abrupt effeotimness lossee experienaed by the 20-percent-ahord flap at superoritical apeeds, huwever, this flap contirmee t o Fram an aaalysia of the lift-cmn%rol dmraoteristios of 8 IO-proent-ohord plain flap on t h e =A 65=210 afrfofl and f r c a n a oompariscm o f =e effeotianeee of this devi08 with that o f a 20-psroent-ohord flap, also emnplaged op the X M A 654210 seotion, the fol’Lopping oonolueions are indioatedt I . For a g i m defleofiOn8 the W% inorement produced by the lO-pemnt-ehord plafn flap at Hemh numbers up to Or876 is never less than 50 peroent of i t a 1 ” e d wlue.
2. Although the effeoWvenesa loasss at superoritioal glpeeda me considerably less aewre for the lO-peroent-chord flap than for the 20-peraent-ehord flap, the larger-ohord flap retains greater effectiveness throughout the Ma& mniber range * o m 0.8 to at least; (3.8750 6 NACA RM No0 'A8A21 TABLE I ORDINATES FOR THE M C A 66-210 A I R F O I L SECTIOH [Statione and ordinatesgiron in peroent of airfoil ohordl
T
Station e566
. 822
1mS31 2 &92 6 . I 0 2 7.606 IO0106 1s 0 101 200091 2s 0079 300064 35 . a 9 40.052 46,016 ,6OoOOO 6 4 . 9 8 6 5 9 . 9 7 3 6 4 . 9 6 4 69 0967 740956 790956 840962 89 e972 940986 1 0 0 . 0 0 0 LoEo radiuar 0 . 6 8 7 ilope of radius through L . E . : 0.0842 Enlarged view of flap f i g u r e f r Typid profile of the NACA 65-210 airfoil secfion with IO-percenf-chord plain flaps o f true airfoil contour.
8 NACA R h ! NO. A 8 A Z 1
. /8
.
./6 . / 4 Mach number, M (0) 6f = -//.8* figure 2 .-The voriafion of secfion drag coefficient wifh Mach number for fhe NACA 6.5-210 ofrfoi/ wifh o 10-percenf-chord plain flop.
. / 6 .14 .04 .02 .6 .7, .8 .9 .3 .4 .5 " 0 " .3 . 4 .5 .6 .7 .8 .9 Mach number, M 12 NACA FUd N O . A8A21 " 0 .3 .4 .5 .6 . 7 .8 .9 LO Mach number, M (e) G f =4.0" Figure 2 r Gonfinued. NAGA 65-210 airfoi/ with a /O-percenf-chord p/ain flop.
c (g) 6, = 1 0 . 6 O Figure 2 : Concluded. NAGA 65-210 airfo// with a lo-percent-chord plain flap.
.
16' NACA R X NO. A8A21
Figure 3 .- Continued NACA 65-210 airfoil with a /O-percenf -chord
plain flap, 18 NACA RM NO. A8A21 ..
(dl 6f'/.7"
Flgure 3 .- Confinued, NACA 65-210 airfoil with a IO-percenr- chord
plain flap.
HACA FIB NO* A81521 19 NACA RM R o ~ A8A21 .
NACA RM NO. A81121 21 -16 -12 -8 -4 0 4 8 1 2 1 6 -16 -I2 -8 -4 0 4 8 I2 1 6 Flq deflection, 6f, deg. nap defle&m,, gf, deg.
5 2
Figure 4.-The wuriition of increment in section lift coefficient with flap deflection at various Mach numbers 21 for the NACA 65-210 airfoil wifh a IO-percent-chord plain flap.
I 1 I I c -I2 *8 -4 0 4 8 I2 I6 H a p deflecfion, 6f, deg.
Iu W I .
(b) for 9 =0.2, Gf =O s
r3 P a s Figure5 .-The variafion of increment in sectfon lift coefficient with Mach number af various flap deflections ~ for fhe NACA 65-210 airfoil with a IO-percent-chord pfahflap.
1 L .- 0 .3 .4 .5 . 6 .? .8 .9 L O M& nmber, M figure 5.-Concfuded. NACA 65-210 airfoil with a IO-percent-chordplainflap.
Mach nmber, M Figure 6.- The variution of flap effectiveness wifh Mach number at various /iff coefficients for the NACA 65-210 airfoi1 with a IO-percent-chord pluh flap.
Mach number, M Figure 7. -Comparison of thevariationin flap effectiveness wifh Mach number at lift coefficients of 0 ond 0.2 for the NACA 65-210 airfoil wifh /O-percent- and 20- percent -chord plain flaps.
.