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.—.-— --- A -..-, 23Y lfilton 3. Ames, Jr., and Richard 1. Sears Langley Afemorial Aeronautical Laboratory e- w.r.i ,, .- . . . . ...
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1.- NATIONAL ADVISORY COMMITT_IIE FOR AERONAUTICS \ L .7 ---—— ---- . , TECHNICAL NOTE NO. ‘?’59 v -—--.-—.
“PRESSURE-DISTRIBUTION INVESTIGATION OF AN N,A,C,A. 0009 .- t AIRFOIL WITH A 30-PERCENT-CHORD PLAIN FLAP AND THREE TABS , .
SUMMARY Pressure-distribution tests of an N.A.C.A. O?09 air- foil with a 3!3-percent-chord plain flap and three plain “+, having chords 10, tabs, 20, and 30 percent of the flap ‘ chord, were made in the i’?,A.C.A. 4- 3Y 6-foot vertical tunnel. The purpose of these tests was to continue an in- vestigation to supply structural and aerodynamic section data that may be ’applied to the design of horizontal and vertical tail surfaces.
The results are presented as diagrams of resultant “pressures and of resultant-yressure increments for the airfoil with the flap and the 25-percent-ckord tab, In- .
crements of normal-for’ce and hin~e-moment coefficients for the airfoil, the flap, and the three. tabs are S.lSO given.
At all unstalled fiap and tah deflections, the exper- ‘ % imental distributions agree well with those calculated by an analytical method. The agreement is poor, however, when the stalled or the unstalled condition of the flap .
or the tab deflected alone was changed to an unstalled or stallqd condition by the simultaneous deflection of both the flap and the tab.
INTRODUCTION .
.
The trailing-edge tab has proved to be an effective ...—.
l “ device in reducing the excessive control forces res’~ltitig from the recent increasesin size and speed of airplanes.
a’ ~ Although s, number of investigations have bepn conducted to determine the characteristics of the different factors N, A. C.A. Technical Note No, 759 ,.- .“ affecting control surfaces (references 1, 2, and %), no data giving the aerodynamic section characteristics of a thin airfoil as affected by flaps and tabs seen.ed to he available that would he applicable to tail-eurface desi:;n.
An investigation was therefore undertaken to supply in- formation applicable to the aerodynamic and the structural design of tall surfaces with ta%s.
The first pert of this investigation comprised pressure-distribution tests made of an N.A,C,A. 0009 airfoil with a 50-percent-chord plain flap and three plain tabs; the, results o: these tests arc reported in reference 4.
The results reported herein were obtained from pres- sure-distribution determinations over one section of an N.A,C,A. 0009 airfoil with a 3!3-percent-chord plain flap and with plain t~”os 10, 20, and 39 percent of the flap chord. From the data obtained, normal-force and pitchin#- moment coefficients were calculated for the airfoil sec- tion complete with the.flap ,dud..the various tabs.
The normal-force and the hinge-moment coefficients for the flap withthe different tabs and.for the tabs separately were also determined. “ A~pA~A~~fj AND T~s.Tcj Model and. Test Installation # The tests vvere made in the. N.A.C.A, vertical wind tun- nel. The test section of this tunnel has been converted from the original open, circular, 5-foot-diameter jet (ref- erence 5) to a closed, rectangular, 4.- Yy ~-foot throat shown in fi~ure 1.
The rectangular 3-foot-chord by 4-foot-s”pan model w,as made of laminated mahogany to the N,A..C.A. 0009 profile.
It was equipped with a plain flap having a chord 30 per- cent of the airfoil chord, c, and with three serially hinyed plain tabs having chords 13, 2(I, and Zg percent of the flap chord, Cf , as shown in fiqure 2. During tests, all flap and tab gaps were sealed with plasticize and cel- .
lulose tape to prevent air leakaqe at the hic+es.
The b radius of curvature at the hinge for both the flap and the t~.bs was approximately one-half the airfoil thickness at the respective hinge positions.
,, A single chordwise row of pressure orifices was built N. A, C.A,, Technical Note No. 759 into the upper and the lower surfaces of the aiifoil, the flap, and the ta%s at the midspan, The orifice positions are shown in fiqure 3, is the model completely spanned the test section, The model was at- two-dimensional flow was approximated, tached to the balance frame by means OT to~que tubes, which extended through the sides of the tunnel and kere rotated by a calibrated electric drive. to set the angle of attack, The flap and the ta% anqles were set inside the tunnel by varying the position of small lever arms on the movable surfaces, ., The rubber tubes from the pressure or,ifices were %rought out of the model at one end through the torque tube and the tunnel wall to a photographically recording ..,.
multiple-tu%”e ‘manometer, b “\ Tests T,es~s were “conducted at an effective Rey~olds Ntimber (Effective Reynold5 “Number = of approximately 7,’4-10,000, The turbulence test Rejnolds Number x turbulence factor.
factor of the 4- by 6-foot vertical tunnelis 1.93’.) The tunnel was operated at,an average dynamic pressure of 10.8 pounds per square foot, corresponding to an air speed of about 65 miles,,fier hour at standard sea-level conditions.
For d’irec”tcompa~isonwith the results presented in reference 4,, the tests were made a% angles of attack from The mctiel was tested -14*O to 10*0 at intervals of 5°, with the 30-percent-chord plain flap deflected 0°, 5°, 10°, 200, 3(30, and 45:9 Throughout the entire angle-of-attack , range for each flap deflection, the three tabs were de- flected 0°~ ~10°, *20°, and +30°, ,.
,,... , ,,, . . . . . . . ,’ ‘$ RESULTS ,, Presentation of D“ata The results of the distribution of pressures are given in the form of diagrams of resultant pressures and .
result:ant-pressure increments, which represent changes in resultant-pressure distribution caused by a change in an- gle of any one”part Qr any combination of the component .“’ N, A. G.A, Technics.1 ,Note No; 759 parts of the airfoil. The resultant normal pressure at any point along the chord. line of the airfoil was deter- mined by taking the algebraic difference of the pressures normal to the surface of the airfoil at that point. All diagrams of resultant pressures or resultaht-pressure in- crements of the airfoil, flap, and tab combination aro plotted as pressure coefficients, P, or as AP, where P-PO .
-— ——.
P= ,, !I’ static _prqssure at a point on airfoil.
static preesure in free air stream.
dynamic pressure of free air stream.
Resultant-pressure diagrams are given for the basic ?- section (i.ec, flap and tab neutral) in fi%ure 4. The resultant-pressure diagram for any other condition may %e obtained hy adding to the basic diaqram the resu.ltant- pressure-incremelnt diagram (figs. 5 to 10} for the partic- ular condition.
The large quantity of data prohibited the inclusion of all the resultant-pressure-i ncrem.ent dia~rams, Only the diagrams for tab deflections of 0° and *30° for the 0.20cf tab, which was considered to be an average size, are presented- Vdlues of angle of attack were selected to represent the followin< conditions: Unstalled negative angle cf attack . .,.-9+0 Low positive angle of attack .. ... ... ..
ho Unstalled high angle of attack . . . . . . . . 5~” The angle of 5~” was selected because it was the highest unstalled angle obtained at some of the higher flap and tal! deflections and because the results COuld he compared with those presented in reference 4.
The section characteristics of the airfoil, the flap, and the tab, as functions of flap and tab deflection, are also plotted “as increments.
These increments were obtained by deducting the basic section coefficients from those for N, A. C,A, Technical Note 3T00 759 5 the “section with the tab, the flap, or the combination de- flected. The characteristics were oltained in each case by mechanical integration of the original plotted pres- sure diagrams.
Computations were made to determine the section coef- ficients, which are defined as follows:
c~ = –n-
airfoil section normal-force coefficient.
qc airfoil section pitching-moment coefficient alout quarter -chord point of airfoil.
nf flap section normal-force coefficient.
Cnf = ‘–- qcf .4P, . .
.hf chf = ‘—~ flap “section hin%e-moment coefficient, .
qcf m ‘t = --- ta% section normal-force coefficient, Cnt qct ht Ch = –—~ tab section hinge-moment coefficient.
t qct wher”e the forces” and moments per unit span are : p normal force of airfoil section.
m pitching moment of airfoil section a%out the -quarter-chord point, nf normal force of flap section.
hf hinge moment of flap section, ~ nt normal force of tall section.
.ht hin%e moment of tab section.
.
and c chord of la,sic airfoil with flap and ta% neutral.
* Cf flap chord, Ct tab chord.
‘6 N. A.. C.A. Technical Note .lio..759 u angle ‘of attack.
a &- 8 flap or ta% deflection.
The sulscript f refers to the flap with the ta%; and the subscript t, to the tab alone.
The integrated coefficients for the basic airfoil are The incre- plotted against anqle of attack in figure 11.
ments for various tab and flap deflections are presented in figures 12 to 20.
Precision # Inasmuch as no air-flow alinement tests have been *’ made in this tunnel, the absolute value of the angle of attack is not known. An error of 1/2° in alinement ap- .
peared t’o have exieted (see reference 4); corrections for ~.
this misalinement were made in the final data. Relative Absolute angles of attack are accurate to within @lo.
flap and tab deflections were set to +2°; their relative settings are accurate to within +lO.
Plotted pressures are correct to within *2 per”cent exc-ept at the yeaks that occurred at the hinge axes and at the nose, The dy- where the variation may be %reater.
namic-pressure readings are accurate to within *1 percent, Two-dfmensio.rial flow having leen approximated, the results may be considered as section characteristics ex- cept for the tu”n.nel-restriction corrections which were applied only to the airfoil section normal-force coeffi- cient, cn. Although no corrections were made for the other coefficients-, they are’ believed to be higher than the free-air values and, hence, are on the conservative side for structural p’urposes.
The magnitude of tunnel corrections for flap and tab coefficients h,as not been de- termined.
The magnitude of the airfoil normal-force co- efficient as represented in the resultant-prsssure- increment diagrams -(figs, 5 to 10) is kaown to be too large by about 11 percent because these curves were plot- .
u . ted directly from tunnel datzi “~PithOut applying any cor- rection.
N,A,.C.A, Technical Note No. 759 ‘7 DISCUSSION Pressure Distribution The distribution of resultant pressure for the basic N.A.C!.A. 0009 airfoil with flap and tabs neutral (fig. 4) and the distribution over the basic section of increments of resultant pressure caused by “flap and tab deflections Such (figs. 5 to 10) are uncorrected for tunnel ei’feet.
diagrams should prove useful in determining loading con- ditions for the structural design of %oth horizontal and vertical tail surfaces.
For this purpose: all pressures are conservative.
This conclusion is especially true of the peaks at the hinge axes because of the use of sealed . gaps.
.* Tests have indicated that the increments of pressure distrilmti,on and the increments of section aerodynamic .k- Coefficients due to flap def].ection are approximately in- dependent of the basic section for conventional airfoils of the same maximum thickness.
It is therefore believed that , for structural design, the incremental data present- ed in this report may be applied to other basic sections of a conventional shape and the same thickness.
Deflection of the flap alone or the tah alone causes an increment of -pressure over the entire airfoil, this ,in- crement reaching peak values both at the nose and at the hinge axis. The increment tapers .from this peak value at the hinge axis to zero at the trailinq ed$e.. When the flap .
and the tah are simultaneously deflected in the same di- rection, the ,increments are largest; peaks occur at the ,..
same places a,nd reach their maximum values.. “If the flap and the tab are deflected simultaneously but in opposite directions, increments are a minimum; but. peaks still OC-’ cur at the nose and at the hinge taxes. The peaks at the” hinge axes are, however, The resultant- of opposite sign..
pressure increment on the flap (flap deflected downward) is a positive peak at the flap hin?e a,xis, passes through zero between the fla”p and, the tab hinges, and reaches a negative peak at the ta% ,axis (tab deflected upwanti), from which it drdps to zero at the tr~,iling edge of the. air- foil.
G In general, the curves of the resultmnt-pressure in- crements are similar in form to those for the 50-percent-’ chord flap of reference 4~, As miqht be expected, the 30-.
N, A., C.A. TecHnical lfote Xo.. .759 percent-chord flap stalls at s,.hiqher angle of attack and its loads are smaller.
The irregularities in some of thp curves (figs.
5(a), 5(c), 6(c), 7(c), 3(%), ~<c), $?(c),:~nd 1O(C)) over the leading-edge portion of the nirfoil may be due to laminar separation caused by severe adv~rse pressure gradients.
.
The curves” of resultant-nressure distribution over the basic airfoil (fi#.
4) ar~ the same as those presented in fig.~re 4 of referen~e 4, The comparison of the experi- mental curves with those obtained by a computed method are therefore omltted”in this report.
Calculations were made for the chordwise” distribution of pres”sure i~crements, resulting from flap and tab dei’lec- .
tions, by the method advanced in reference 6.
Comparisons .
with the experimental curves are shown in fisures 5, 7, 8, and 9, The curves in these figures show this method of computing incremental pressure distributions to be in ‘b good’ a+reement with experimental results for deflections of the c~s~s at which the flap or tab alone W2,S unstalled.
for which agreement ~iay. be..considered satisfactory , the greatest divergence of the curves is 0.2AP and occurs at the hinge axes (fiqs. 5(a) and 7(b)), ~~~en the flap and the tab ~~ere sim- On the other hand, ultaneously deflected in the same direction or in opFo- site directions and the flap, the tab, or both were stalle&, the computed curves did not agree with the ex- perimental curves.
This result i’s in disagreement with other results cited i.n reference 6.
.
The method advanced in reference 6 is based upon the ..
assumption that, at a qiien angle of attack, the coeffi- cient increments Acn and Acm for the flaps deflected alone to a given anq’le ~lus the coefficient increments for the tal deflected alone to a given angle should equal the coefficient increments for fla~ and tal simultaneously de- flected to these qiven angles,- This assumption is not borne out by experiment in the folloming cases: Flap deflection, 10°; anqle Of attack; 5~”; .
d.
tab deflection, -30° (fig. 7(c)).
3“lsLpdeflection, 200; angle of attack, 1/2°: tab deflection, -30° (“fig. 8(b)).
Flap deflection,” 300; zmql”e of attack, @o; tab deflection, *30° (fi#. 9(b)).
N. A. C.A. Technical Note No. ’759 9 This poor summation of coefficients is caused %y the occurrence of a critical condition in which the stalled or the unstalled condition of either the flap or the tab deflected alone is chanqed by the simultaneous deflection .
of %o.th the flap and the ta%, The experiments show that., in some cases where t,he” flap was stalled when deflected alone , the tab when deflected at the same time in the op- posite direction caused the flap,to %ecome unstalled. In other tests, where both the flap and the tab were deflect- ed in the same direction at the same time, the ta% often became stalled and, in a num%er of cases, both the flap and the tab stalled. It is evident that the method of computing chordwise distribution, %ased on the summation of coefficients due to the flap and the tab deflecti~ns, will not check the experimental results where the stalled or the unstalled condition of the flay or the ta% deflect-= k ed alone was changed by the simultaneous deflection of the l flap and ta%, It should be noted here, however, that var- . .
ious free-flight tests have shown that,” for these critical conditions, the &talls in free flight may not necessarily A“,” occur in the order that the tunnel tests have indicated.” ., Aerodynamic Section Characteristics Airfoil characteristi~~.- The.basic airfoil section .,.—— _______ _______ ______ __ gave a linear variation of Cn a%ainst angle of attack .(-fig, 11) in the unstalled range, w~nich was similar to the results ol)tained in reference 4.
The slope of the norma.l- force curve, is 0,095 and agrees with the results Zcn/az, in references 4 and 70 The failure 0$ the en curves to .
give ,a value of Cn = O at a = O was probally due to model imperfections and tab r.isal.inement.
The maximum in- ,.
crement of normal-force coefficient, Acn = 1.73, occurred at a = -9$0, fif = 45°, &n$. et = 30° for the 0,30cf ta%o (See f’i<. 18(3).’) When the tab was deflected upward to -~oo , the other conditions remaining the same, the value of Len %ecame. 1.06.
The reduction of hen “due to re- versing th+. ta% WaS ().67, or 39 percent, For the same condition in reference 4 where the model had & 0.50c flnp, the maximum Acn was 2.32 for the ta% deflection of 30° and 1.34 fc,r the , -309 t,n.b deflection. The chari+jein Acn . .
in this case WR,S !3.98, or ,alout 42 percent for & 0.50 flap.
In fiqures 12, Is, ,Pund18, the curves change slope Acn .
at a flap deflection of 200 for negative angles of attack .
and between flap deflections of 10o and 20° for the posi- tive angles of attack.
These changes in slope are proha- ‘Dly caused hy the stalling of the flnp~ 10 N, A. C.A. Technical 3Tote No. 759 The pitchinq-moment coefficient for the br, sic n.i. l’- foil section was anmroximately zero for the range of angles of attack o;--9”~0 to 10~0, (See fig. 11, ) The value of varied nearly linearly with flap deflect- cm ion for all values of m within the unctalled range.
It was noted that the effectiveness of the tabs in tiar;:- decreased as tab deflections increased posi- ing ACm tively or negatively from the neutral position.
Flare and tab characteristfcs,- In agreement with ref- .-— --————--___-—.-_.——-—-— erence 4, the increments of flap section normal-force an?.
hinge-moment coefficients varied nearly linenrly with flnp deflection within the unstalled range of the fla;p. As would be expected, the flap stalled at succ~ssively lower flap deflections as the angle of attack we.s increased.
.
The curves of AcnP and Achf were shifted para~- .
lel to themselves with different tab deflections. The rate of change of the increments with tab deflection de- -b.
creased as the tab deflection increased positively or negatively from neutral, ~0° (upward] In most cases, the -.
deflection of the tab was rather ineffective in reducing as shown. by the irregularity of the increments for %f r this tah deflection.
This result aqrees with the results in references 1 and 2.
The incremental tah section coefficients, Acnt and plotted in fiqures 14, 17, and 20, varied linearly @t ~ with tab deflection through the unstalled rarige and were .
larger when the flap and the tah mere deflected in the same direction than when they were deflected in opposite directions.
.
As indicated in the curves showin$ the flap-coeffi- cient increments, the tab deflected -30° was stalled in most cases. In aqreement with reference 4, at given va~- ues of a and an increase in flap deflection caused at , increases in Acnt and Acht. As the flap deflection was increased, the magnitude of the increases in Acnt and .
ACht generally became larger. (See fiqs. 24, 1’7, and .
20.)
.
b N, A. C.A. Technical Note No. . 759 The angle of attack for the curve of the flap de= fleeted 300 in figure 14(f) is 9° instead of l@O as for Since the model in this con- the other flap deflections.
dition was partly stalled, the flow was too unsteady for data to he taken at a = lo~” l . .
Comparison with Other Tests In the following talle are listed some of the more important average slopes o%tained from this investiga- tion. The comparisons are made with the experimental re- All slopes are sults from data in references 3 and 4.
corrected to infinite aspect ratio.
#--A-
,.
.
..
...
,4 ,.
,..
.
lJ.A. g.A. 0009 AIRFOIL ——. .—— —.
Effective Flap chord &hf G&j
Source Tunnel Reynolds :percent c) P
Nuder
-0.52 -0.014 ‘-o .013 ao.;l
4- by 30 0.095 0.57 3,410,000 This &foot report i- -.0096 b.25 “-.013 ‘8 .0 -.70 l?lll- 41 1,606,000 l OE3 Refer- .
scale ence 2 a.29 4- “by 3,41O,OOC .055 l 86 -.76 -.016 9-.016
Eefer- 50
ence < 6-foot a Tab size, 0.20cf.
b “: Ta~ size, o.195cf.
Ct CD s , a -.
w r . >’
F’ l
r N, A. C.A. Technical Note No. 759 * Two notable differences occur in this comparison be- -4 tween data from tests in the 4- by 6--foot vertical tunnel” and in the full-scale tunnels the value of acn/aa First, of 0.095 obtained from the ’present tests and those reported in reference 4, although not in accord with the value of 0.083 obtained in the tests reported in refere~.ce 3, is in agreement with the value of 0,095 from ‘other full-scale- tunnel tests reported in reference 7.
The other differ- ence occurs in comparisons of values of 3chf/86f and This difference is in accordance with expected ~ch+~~.
results. As pointed out previously, the section charac- ‘ teristics of the flap and the tab herein reported were not corrected for tunnel effects.
In addition, the gaps be- tween the flap and the tab were sealed. Both of these factors would cause an increase in these slopes. .
\ * >.
CONCLUDING REMARKS. .’ ,., Aerodynamic section characteri,stics,and resultant- pressure dist~ibutions have %een presented for the N.A.C,A.
0009 airfoil ,with a Go-percent-chord flap and three plain tabs havi~ng chords 10, 20, and 30 percent of the”,flap chord.
For all’ unstalled flap and ta% deflections, the ex- p~rimental,a,n,d the calculated distributions of resultan’t- pressure increments are in good agreement.
.
.
The results of the analytical method of calculatin~ the resultant-pressure distribution apd the ‘experimental . .
results are not in agreement for the cases ~ri which the stalled or the unstal~ed condition of the flap or the tab deflected. alone was changed by the simultaneous deflec- tion of the flap and the t’at.
This poor agreement between the experi,mential an~;the calculated results,is attributed to the fact that the coefficient increments for these critical conditions are not additive, ,as they must be to obtain good agreement.
.
In the application of these data for design purposes, .
it should he remembered that, for all cases, gaps were completely sealed, resulting in higher peak pressures at the hinge axes and in higher hinge-moment and normal-force coefficients than WOUlti have %een obtained with unsealed .
14 N. A.. C.A. Technical Note No. 759 l gaps. It should also be noted that only the values of the F- normal-force coefficients were corrected for tunnel effects, Langley Memorial Aeronautical Laboratory, National Advisory Committee for Aeronautics, Langley Field, Vs., March 27’, 1940.
REFERENCES 1.
Wenzinger, Carl J,: Pressure Distrilmtion over an T..R, NO. 574, Airfoil Section with a Flap and Tab.
N.A,C.A. , 1936.
.
* 2.
Earris, Thomas A.: Reduction of Hinge Moments of Air- plane Control Surfaces bay Tabs. T.R, No, 528, N,A.C.A. , 193!5, “A 3. Goett, Effects of Eleva- Harry J., and Reeder, J. P, : tor Nose Shape, Gap, Balance, and Ta.%s on the Aero- dynamic Characteristics of a Horizontal. Tail Sur- T.R. No. 675, N.A.C.A. , 1939.
face.
4. Street, William G., and Ames, Milton B., Jr.: Pressure- Distribution Investigation of an N.A.C.A. 0009 Air- , foil with a 50-Percent-Chord Plain Flap and Three Tabs, T.N. No. 734, N.A.C.A., 1939.
.
5. Wenzinger, Carl J., and Harris, Thomas A.: The Verti- cal Wind Tunnel of the National Advisory Committee for Aeronautics.
T.R. No, 587, K.A.C,A., 1931, ., 6, Allen, H. Julian: Calculation of the Chordwise Load Distribution over Airfoil Sections with Plain, Split, or Serially Hinged Trailing-Edge Flaps.
T*R* No. 634, N.A.C.A. , 1938.
Croett, Harry J., and Bullivant, W. Kenneth: Tests of N.A,C.A, 0009, 0012, and 0018 Airfoils in the Full- .
Scale Tunnel.
T.R. No, 647, N,A.C.A. , 1938.
.
.
, .
F~B. 1.2,3 B. A.C.A. ?oohic81 SOte ik. 759
(~
-(.S
q
‘1~
Is
........--------
& of bo/ante
b
.- -- —, —-— —K ---- sys7’em ---7
-------- -------
Reci%ngular fest section.
-.’
I
II I Model mmmted im 4-by 6-foot rertioal tumel.
F*o l.-
--F-
.Q of orifices ~ <-- .
..................------- -----------------------
$
%
A:A .
..---- ,, with 0.30c plain flap Figure 2.- The H.A.C.A. 0009 pressure-dintribution model and O.lOcfj 0.20cf,j l nd 0.30cf pluin tcbs.
,.
m- a
}ri- Lo- 1’7 ‘70.0 Fice ca,- 18 72.5 tion 75.0 :: 80.0
77 1 .62
21 85.0 1.25 22 87.5 2.5 23 90.5 5.0 Q) w 24 91.0 5 10.0 .
$ 25 91.5 6 20.0 & 92.5 7 30.0 % 93.5 40.0 28 94.0 9 45.0
$QJ 94.5
10 ]48.2 I
‘v :: 95.5 11 150.0I 31 96.5 i2 52.5 orificoa en the NtA,C.A.
Figure 3.- Chordwiae locations of pr.ssurc 32 97*O
13 55.0 I
0009 airfoil in percent ohord.
97.5 14. 60.0 34 98.5 15 65.0
I
99.5 — ....
r.
Note No.759 N.A.C.A.Technical Fig.4 ;, F
+
ID====
c- ( ‘- ~
l / .
. .
-2
—. .-.-
a=’-’1f2~ ~ (
—
-4
A“
I ——--—.
-6
--t
/ !
-.- ——.— .--—— .—. —,..
—.- --J P 1/20 P o I .
,.
\
—- —— ——
--
5 1/2° ,
‘“ \. (
—--- ———.—.-——. -- —— ——
. m
- __ ‘$ .
20 60
80 1 ,00
Percent chord Distribution of resultant pressure over the N.A.C.A. 0009 airfoil at various angles of attack: Flap and tab neutral.
i ,’ ti “* .
T 4-, L ?
*’C + ( --Q — AP \. ~ \u~ I \ 6“ - 3b” I I Iflbi I -.
“.
u 1 \ v 1 1 1 I I I o .
I I I I 1/ I I I I -20 100 0 20 40 60 80 /00 o 20 40 60 80 loo 20 40 60 80 Percenf chord Percemf chord Percenf chord Figure” 10.-Increments of resultant pres-.~ Figure 8.- Increments of resultant pres- Figure 9. -Increments of resultant pres- sures for various angles of attack and sures for various angles of attack and .0 sures for various angles of attack and various deflections of a 0.20cf tab on z various deflections of a 0.20cf tab on a various deflections of a 0.20cf tab on 0.30c flap deflected 200. a 0.30c flap deflected 45Q.
a 0.30c flap deflected 300.
I f’ .. -k’ , * ,’ , l — 0.70C, A––--–– .20.
a—— .30 “ + <? ~ + .
— - .6= --~ ., \ /, -u H : -a- . == .
~ -- ,.
/ 7 6/ 1’ d ; -8 0 8 AngJe of offock, d, deg Fee 11. - Section characteristics of basic N.A.C.A..0009 airfoil with 0.30c plain flap and tabs,neutral. ~ w I f II ,, k Fig. 12a,b,c N.A.C.A. Technical Note No. 759 .
,; ,’n- - .i 1 ., T — ., .+ .. ..
.:i .- —.
.
-.
.. l -.
.— l .
.
.-— -4 1/20 (b) a= -9 1/20 (c) u= (a) c,= -14 1/2° Figure 12, a to f.- Increments of airfoil section normal-force and pitching-moment coefficients for varioue deflections of a 0.30c plain flap and a O.10cf tab.
- .* Fig. 12d,e,f N.&C.A. Technical Rote No. 759 -.
-- . .S . . . .
.
t — .
s ..> . .- .-.
.
.
-.
.
.
.. >-.-U Figure 12 concluded.
Jncremen+ of flap sec +ion normol - force “coefficient Ac~, % ,4 &c:G&gb a“&Aika”N Abb CY ‘3 ~ Q1 J o ?$ s -0: $+8 ~ /ncremen+ of flap section hiwe-momen+ coefftkient Ac,.
nn . .
.
Incremen+ of flap section normal-force coefficient, At.{ * ..
.
.
l ’S.
.- --=. ----- P., # , .“ .
, .4 1.8 I 1 1 I I 1 I I 1 I I I I I I I I I I I I .3
II
/.6
H- E13=E%Ii
.2 J!
c1 :..1 .~ ~o $ $1 &z ~“ E I –.3 al cm c ~ ..4 c o jj -.5 $’ s-.6 a & -.7 $_a ~“ b-g <“ -1.I -1.c -1.2 ,.
-1.2 (c) o 10 20 30 40 50 10 20 30 40 50 Flap deflection, C$f. deg a = -4 1/20 El Figure 13 eontinned.
.0 9.
Incremenf of fbo sec+ion normal-force coefficient. A c..
r .
l II m l -.
.
.
* “1 .
.
, .
.
1.8- 1.6 J ::mll%;!efi [1, /.4 t 1 .
~- 1--.1 --I. -1---1 --+ --!
:.3
~’/. o
/ < .. ..
— / W+l--l--l-;--;-l; /’ j ..-, . .. ....++.
$.2 * .8 &i /1 : ? ‘% .
~ E QJ g -.8 ~ -.6 < $ -.7 -Lo :8 -1.2 -.9 -/.4 -30 -20 -10 0 /0 2’0 30 W5v -20 -/0 o 10 20 30 Tab deflection, 6, , deg Tab deflection, 6, , deg a= -4 1/20 a = 1/20 /@*& Figure 14 . . m~inued.
,, ,: !’; ”, .,,1.1 I , . .
I I I I I I I I I I I I O-—-—- 2O-.LI I I w --4 / / ,$ .3 &.o ), .: ..- / .. ..
/1 I lJ L1 !. I I I I I I : I .M$ L-u+H+H+l ~ # _ / / -:.
‘.0 I I I I I I x. I I II <-.4 %.
% ; -i. .
~ 75 -i. -.5 ~ $ 5 E ~ _.6 c QJ -, ~ -.6 ? ~ b & ~ ~ & -.7 -1 -.7 -.8 -.8 -1.
-.9 -.9 -1.
-30 -20 -/0 o 10 20 30@%o -20 -/0 o 10 20 30 -30 -20 -10 0 /0 20 30 (f )-30 -20 -lo 0 10 20 30 Tab deflection, 6, , deg Tffb deflection, C$t , deg ~ . 10 1/20 n,= 5 1/20 -Y u.
Figure :!
I . 4 z “* .
.
.
1- Cn .
Incremenf of airfoil Se&I-On normal-force coefficient, A cm ,% -$ ‘a .
Fig. 15d,e,f N.A.C.A. Technical Note No. 759 . .
* .
c “.
.
.
Flqodef/ecfion, G,deg (e) U= 5 1/2° (d) u = 1/2° (f) u = 10 1/20 Figure 15 concluded.
* .
* .7 .6 .5 ‘.4 .3 .2 , , I I , t I Flop deflection, 6J, deg Flop deflection, 6$, deg .9 ~o a= a = -14 1/2° Fi~ro 16, a tof. - Increments offlap Soctien normal-force and hinge-rnommt ooeff icients for variouu def lacti~e of a 0.300 plain flap and a 0.20cf tab.
,, , u .
$
$2
~--6 ~ ‘“ >.
?
c x =_ -.7 .4 -.7 -.8 -.8 -.6 I I I I -.9 -.8 -.8 -.9
El
(c) o /0 @J o /0 20 30 40 50 10 20 30 40 50 o 20 30 40 ! io 0 10 20 30 40 50 Flap def/ecfion. 61, deg flap deflection, 6,, deg w Figure I I I I -3 -.8 II IL II II III Ill (f) : ‘ /! 20 30 40 50 o 10 20 30 40 50 flap deflection, 6“, deg flop deflection, 6’, deg w 1$ ~ . 10 1/20 a = 5 1/2° z Figuro 16 comluded.
.0 1+ 1, * . , b, .
% az +ij-.r?l i, i i i-,j I I x I WI \ I %.
o % -.4 -.-6 ~: 4 $ $“ ~ E & E ?-.5 y -.5 F-.8 g & < : $ -.6 -Lo -.6 -.7 -/.2 -.7 -.8 -.8 -L4 [ 20 30 @’)-30 -20 -10 0 10 20 30 -30 -20 -lo 0 /0 20 30 (@-30 -20 -10 0 /0 20 -30 -20 -lo 0 10 Tab deflection, dt, deg Tab deffecfion, &t, deg a= -9 1/20 a= -14 1/20 $ Figara 17, a to f.- Incromrnta of tab saction normal-force and hinge-mmmt caefficiento for mriow deflections of a 0.300 plain flap and a 0.20cf tab.
z * .
u Increment of fob section normal-force coefficient, At..
.
,* .
.
— — .
-.
.— .— Incremenf of +ab section normal-force coefficient, Ac..
Incremenf of fob sec+ion normal-force coefficient, A c..
* .
.
-.
Incremen+ of fob section normol- force coefficient, Ac., .
# .
# s .2 N.A.C.A, Technical Note No. 759 Fig. 18a,b,c ...-=.
I , 6*>#Jg_ — 2.0 .2 I — n—.
& 1 A— — 20 — I YW I I ;---------73 I I I
-’EENKE=im
— ..— -.2 .4 — 0 -3 UE i — J -.4 Q ‘“
1 —
:, S.# l — .
.$ b -- .8 (c) -. I .4 — -.2 2.3 -4 J/l — -.4 .
-.8 ~ ,0 Z. so do so /0 20 30 4’0 50 o Flop deflection, dx,deg (b) a= -9 i/20 (c) a.= -4 1/20 (a) u= -14 1/2’3 Lciente Figure 18, a to f.- Incremente of airfoil section normal-force and pitchiw-moment coeffj for various deflection of l 0.30c plain flap and a 0.300f tab.
., H. A.C.A. Technical Note No. 759 Fig. 18d,e,f .=..= > . ..
\ .- . . ,F ..- .“ .-.
.— . ..— — .
..
-.
.
b .
Flop deflection, 6“ , deg (d) a = 1/20 (e) u= 5 1/2° (f) a= 10 1/20 Figure 18 consluded.
, 7 . ..
a flap defi’ecflon, C$f, deg % l-.
~. .14 1/20 Cp 1- s’ Figaro 19, a ko f.- Increments Of flap section normal-force and hinge-moment coefficients for various deflections of a O.?JOCplain flap and a 0.30cf tab.
“u ill: ,.
c * .
.
* /0 0 20 30 40 50 Flap def/ecfion, df, deg.
flap deflection, &f, deg = . .4 1/20 a = 1/20 Figure 19 continued.
I I i, 0 10 20 30 40 .50 Flop deflection, 6!, deg Flop deflection, 6~, deg a= 51/20 a= 10 lf20 Figure 9 concluded..
, I ?
,“, ;’ ,,, ,, iii “1 I ‘11.,:1: I :,11 ,. ,1 Incremenf of iob sec}ion normal-force coe fficien+, Ac., “ —.
_...
Increment of fab section normal-force coefficient, Ac..
d.)
Q $ < Q Q —.
s
IN o —. ~ Incremen + of fob secfion hinge-moment coeff7cienf, ACti.
.
/.8 .6 I I 6~, Aeg – L6 ~ 0+- .5 1.4
“lztEHiE=kH
.
.- 44 C.? 1.2 : t—--’----” a .3 1’ 1.0 .y $4 -.7 -1.0 -n -.8 -1.2 -.9 -.9
I Iim
-1.4 -1.4 30 (CL3O -20 -/0 o 20 30 @)-30 -20 -/0 o /0 -30 -?0 -10 0 10 20 -30 -20 -10 0 10 20 30 10 20 30 -.
Tab deflection,” dt , deg Tab defl=fion, dt, deg ~ . -4 ~/20 1/20 a.= y~y~+ : g Figure 20 continued.
-m ,.
“.
,,, -.8 /.0 -30 -20 -10 0 10 20 30 --- ,“L “u” Job deflection, dt, deg a.= 5 1/20 “m= 10 1/20 Figure 20 concluded.
,:.
I :’ .,