Document
NATIONAL ADVISORY COMMITTEE
FOR AERONAUTICS i_l_.'_ _S
REPORT No. 666 WIND-TUNNEL INVESTIGATION OF N. A. C. A, 23012,
23021, AND 23030 AIRFOILS WITH VARIOUS
SIZES OF SPLIT FLAP
By CARL J. WENZINGER and THOMAS A. HARRIS _EPR DUCED8Y }_ONAL TECHNICAL N,_ .... "-'N SERVICE INFORhAA Ii_ RT_ENT oF COMMERCE U S DEP.A. • VA 22161 SpR|_GFtE_D, " AERONAUTIC SYMBOLS 1. FUNDAMENTAL AND DERIVE D UNITS Metric English Symbol Abbrevia- Unit Abbrevia- Unit tion tion Length ...... l meter .................. m foot (or mile) ........ ft. (or mi.)
Time ........
t second ................. s second (or hour) ....... see. (or hr.)
Force ........ F weight of 1 kilogram ..... kg lb.
weight of 1 pound .....
Power ....... P horsepower (metric) ............... horsepower ..........
hp.
miles per hour ........
kilometers per hour ...... k.p.h. m.p.h.
V Speed .......
meters per second ....... m.p.s. feet per second ........
f.p.s.
2. GENERAL SYMBOLS w, Weight--rag v, Kinematic viscosity g, Standard acceleration of gravity--9.80665 p, Density (mass per unit volume) m/s2 or 32.1740 ft./sec. 2 Standard density of dry air, 0.12497 kg-m'4-s _ at W 15 ° C. and 760 ram; or 0.002378 lb.-ft. -4 sec. z Mass-_- m, Specific weight of "standard" air, 1.2255 kg/m 3 or g L Moment of inertia--ink 2. (Indicate axis of 0.07651 lb./cu, ft.
radius of gyration k by proper subscript.)
Coefficient of viscosity 3. AERODYNAMIC SYMBOLS S, Area • Angle of setting of wings (relative to thrust S_, Area of wing line) G, Gap _t, Angle of stabilizer setting (relative to thrust line) b, Span Resultant moment c, Chord Q, b 2 Resultant angular velocity _, Aspect ratio Vl Reynolds Number, where 1 is a linear dimension V, True air speed I.t (e.g., forL a-model airfoil 3 in. chord, 100 m.p.h, normal pressure at 15 ° C., the cor- q, Dynamic pressure_--_pV responding number is 234,000; or for a model of l0 cm chord, 40 m.p.s., the corresponding L, Lift, absolute coefficient CL_-a_ number is 274,000) C_, Center-of-pressure coefficient (ratio of distance D, Drag, absolute coefficient CD=_ of c.p. from leading edge to chord length) Angle of attack Do, Profile drag, absolute coefficient CD0--a_ 5, Angle of downwash D_, Induced drag, absolute coefficient Cm=a_ 0_02 Angle of attack, infinite aspect ratio Angle of attack, induced Dp, Parasite drag, absolute coefficient CDv--a_ Angle of attack, absolute (measured from zero- O_tt, J- lift position) a}t , Flight-path angle C, Cross-wind force, absolute coefficient Cv=_ R, Resultant force
REPORT No. 668
WIND-TUNNEL INVESTIGATION OF N. A. C. A. 23012,
23021, AND 23030 AIRFOILS WITH VARIOUS
SIZES OF SPLIT FLAP
By CARL J. WENZINGER and THOMAS A. HARRIS Langley Memorial Aeronautical Laboratory 161568--39-----I NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS HEADQUARTERS, NAVY BUILDING, WASHINGTON, D. C.
LABORATORIES, LANGLEY FIELD. VA.
Created bY act of Congress approved March 3, 1915, for the supervision and direction of the scientific study of the problems of flight (U. S. Code, Title 50, Sec. 151). Its membership was increased to 15 by act approved March 2, 1929. The members are appointed by the President, and serve as such without compensation.
JOSEPH S. AMES, Ph.D., Chairman, ROBERT H. HINCKLEY, A. B., Chairman, Civil Aeronautics Authority.
Baltimore, Md.
JEROME C. HUNSAKER, Sc. D., V._..NNEVAR BUSH, So. D., Vice C_irman, Cambridge, Mass, Washington, D. (_.
SYDNEY M. Kl_us, Captain, United States Navy, CHARLES G. ABBOT, So. D., Bureau of Aeronautics, Navy Department.
Secretary, Smithsonian Institution.
CHARLES A. LINDBERGH, LL. D., HENRY H. ARNOW, Major General, United States Army, New York City.
Chief of Air Corps, War Department.
F_cm W. REICHELDERFER, A. S., GEORGE H. BRZTr, Brigadier General, United States Army, Chief, United States Weather Bureau.
Chief Matdriel Division, Air Corps, Wright Field, Dayton, Jos_r H. TowEP.% Rear Admiral, United States Navy, Ohio.
Chief, Bureau of Aeronautics, Navy Department.
LYMAN Jo BRIGGS, Ph.D., EDWARD WARNER, SCo D., Director, National Bureau of Standards.
Greenwich, Conn.
ORVXLLZ WRmHT, So. D., CLINTON M. HESTER, A. B., LL.B., Dayton, Ohio.
Administrator, Civil Aeronautics Authority.
GEORGE W. LEwis, D/rector of Aeronautical Research Joss F. VICTORY, Sectary HENRY J. E. REID, Enoineer..i_l_Charge , Langley Memorial Aeronau$ical Laboratory, Langley Field, Va.
JosN J. InE, Technical Assistant in Europe, Paris, France TECHNICAL COMMITTEES AERODYNAMICS AIRCRAFT STRUCTURES POWER PLANTS FOR AIRCRAFT AIRCRAFT ACCIDENTS AIRCRAFT MATERIALS • INVENTIONS AND DESIGNS Coordination of Research Needs of Military and Civil Aviation Preparation of Research Programs Allocation of Problems Prevention of Duplication Consideration of Inventions LANGLEY MEMORIAL AERONAUTICAL LABORATORY OFFICE OF AERONAUTICAL INTELLIGENCE LANGLEY FIELD. VA. WASHINGTON, Do C.
Unified conduct, for all agencies, of scientific research on the Collection, classification, compilation, and dissemination of fundamental problems of flight. scientific and technical information on aeronautics.
5-24-39
REPORT No. 668
WIND-TUNNEL INVESTIGATION OF N. A. C. A. 23012, 23021, AND 23030 AIRFOILS • WITH VARIOUS SIZES OF SPLIT FLAP By CARL J. W_NZING_R and THOMAS A. HARRIS SUMMARY 230 series were used because they appear to be generally An investigation has been made in the N. A. C. A. satisfactory for most purposes. The high-lift device 7- by lO-Joot wind tunnel off large-chord N. A. C. ,4. investigated with the airfoils of various thicknesses $8012,23021, and 23030 a_rfoils with split flaps 10, 20, 30, was the simple split flap, which is used as a basis of corn- chord from 10 to 40 percent of the wing chord were and 50 percent oj the wing chord to determine the section parison with other high-lift devices. Flaps ranging in aerodynamic characteristics oJ the airfails as affected by each airfoil. These tests are expected to be airfoil thickness, flap chord, and flap de]led_on. The tested on complete section aerodynamic characteristics o] all the foUowed at a later date with tests of slotted flaps on combinations tested are g_ven in the form o.f graphs o] lift, similar airfoils. MODELS drag, and _itching-moment coe_cients, and certain pLAINAIRFOILS applications to aerodynamic design are discussed.
The final maximum lift coe_vients.for Zhe three airfoils Three basic wings, or plain airfoils, were used in these tested with the 020c_ flap were about equal. For the tests; each had a chord of 3 feet and a span of 7 feet.
airfoils with the 0.10cw flap, She ma_dmum lift coe_'w icnt The models were constructed of laminated wood and decreased with airfoil thickness; .for the airfails with the were built to the N. A. C. A. 23012, 23021, and 23030 0.30cw or O.$Oc,r flaps, the mozimum l_fl eoe_cient _n- profiles. The thickness of each of these airfoils is, creased with a_rfo'il thickness to a maximum" vaI_ o.f respectively, 12, 21, and 30 percent of the wing chord, _.95. WitMn the range covered, the increment o.f mazimum c,o. The ordinates for each of the three airfoils are lift coc_dent due to the split flaps was practically inde, listed in table I. The N. A. C. A. 23012 airfoil, which pendent o] Reynolds Number. The increase in minimum had been previously used for the investigation de- profile-drag coe_'wient with a_rfog thickness was large, scribed in reference 1, was already available.
being about twice as great.for the N. A. C. A. 23030 as for FLAPS the _3012 plain a_rfoil. Four simple split flaps extending along the entire INTRODUCTION span were used with each model. The flap chords, or, were 0.10cw, 0.20c_, 0.30c_, and 0.40c_ and were The National Advisory Committee for Aeronautics is believed likely to cover the range of sizes that might undertaking an extensive investigation of various high- be used in practice. (See figs. 1, 2, and 3.) The lift arrangements to furnish information applicable to flaps were built of plywood braced at several points the design of wing combinations for the improvement of along the span and were arranged for setting at de- the safety and the performance of airplanes. Thus far, flections from 0 ° to 105 ° down. The flap deflection, most of the tests have been made with wings having a St, is measured between the lower surface of each air- thickness 12 percent of the wing chord and having the foil and .the flap, as shown in figures 1, 2, and 3.
Clark Y or the N. k. C. A. 23012 profile. It appears TFSTS very desirable at the present time, however, to extend the investigation to include wings having other thick- The models were mounted in the closed test section nesses and also other airfoil profiles. Tim present report of the N. A. C. A. 7- by 10-foot wind tunnel so as to describes the results obtahmd from tests in the 7- by span the jet completely except for snmil clearances at 10-foot wind tunnel of airfoils of various thicknesses each end. (See references 1 and 2.) The main air- equipped with high-lift devices, foil was rigidly attached to the balance frame by The investigation was made of airfoils having thick- torque tubes, which extended through the upper and nesses from 12 to 30 percent of the wing chord; these the lower boundaries of the tunnel. The angle of thicknesses are believed to cover the range likely to be attack of the model was set from outside the tunnel by met with in practice. Airfoil sections of the N. A. C.A. rotating the torque tubes with a calibrated electric • REPORT NO. 668--NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS drive. Approximately two-dimensional flow is ob- An angle-of-attack range from --6 ° to the angle of attack for maximum lift was covered in 2 ° increments tained with this type of installation and the section characteristics of the model under test can be deter- for each test. Lift, drag, and pitching moment were mined. measured at each angle of attack.
A dynamic pressure of 16.37 pounds per square foot RESULTS AND DISCUSSION was maintained for most of the tests, which cor- responds to a velocity of 80 miles per hour under stand- COEFFICIENTS ard atmospheric conditions and to an average test All test results are given in standard section non- Reynolds Number of about 2,190,000. Because of the dimensional coefficient form for the airfoil and flap combinations corrected as explained in reference 1.
FIGURE L--Sectlon of N. A. C. A. 23012 airfoil with split flaps, e;=O.10cw. O.20cw.
0.30c,, and 0.40c,.
cw=36.
/6 Ft(ltrlll 2.--Section of N. A. C. A. 23021 atrfofl with split flaps, e/-0.10c,, 0.2_,, 0.30c., and. OAOc..
, c_=36" 0 .8 .4 .6 .6' 1.0 1.2 1.4 I.B u_ Sec/ion lift coefficienf, c z FIoUx¢ 4.--Section aerodynamic characteristics of N. A. C A. 23012 plain airfoil.
Fl_ua_ 3.--Section of N. A. C. A. 23(}30 airfoil with split flaps, c/=0.10c,, 0.20¢=, el, section lift coefficient, l/qc=.
0.30e., and 0.40e..
cdo, section profile-drag coefficient, do/_c,_.
c_¢_.,.)o, section pitching-moment coefficient about aero- turbulence in the wind tunnel, the effective Reynolds Number, R,, was approximately 3,500,000. For all dynamic center of plain airfoil, m_._.)o/qC_ 2.
where tests, R6 is based on the chord of the airfoil with the 1 is section lift.
flap retracted and on a turbulence factor of 1.6 for the tunnel.
do, section profile drag.
Each airfoil was tested by itself without the flap so section pitching moment.
_(a.c.)O, that the characteristics of the plain airfoils could be _1, dynamic pressure, 1/2 pV2.
determined. Each of the four split flaps was then Cto, chord of basic airfoil with flap fully retracted.
and tested on each of the three airfoils and deflected in 10 ° or 15 ° increments up to the deflection giving the is angle of attack for infinite aspect ratio.
a0 highest value of the maximmn lift coefficient.
flap deflection.
N. A. C. A. 23012, 23021, AND 23030 AIRFOILS WITH SPLIT FLAPS PRECISION The accuracy of the various measurements in the tests is believed to be within the following limits: -t-0.0006 _o ............. 4-0"1° Cd°(c_-l.0) ........
±0.002 c_,._ ........... +0.03 ceoccz.2.5) .......
+0.003 _t ............. +0"2° Cm(a.c.) 0 .........
e_o_ .......... -t-0.0003 SECTION AERODYNAMIC CHARACTERISTICS Plain airfoils.--The section aerodynamic characteris- tics of the N. A. C. A. 23012 plain airfoil, as determined with the two-dimensional-flow installation, are shown I .048 .044 FIGURS 8.--Section aerodynamic characteristics of N. A. C. A. 23030 plain airfoil, coefficient about the aerodynamic center is --0.002 compared with --0.003 for the N. A. C. A. 23021 and --0.000 for the N. A. C. A. 23012. The most marked change is the position of the aerodynamic center of the plain airfoil; it is 11 percent of the _ . 008 chord ahead of the quarter-chord point of the wing and about 44 percent of the chord above the chord .004 line.
_.
b . O • _ 0 .2 .4 .6 .8 LO 1.2 1.4 1.8 i; !: FI_URE &--Section aerodynamic characteidstlcs of N. A. C, A, 23021 plain airfoil.
in figure 4. Similar results for the N. A. C. A. 23021 and the N. A. C. A. 23030 plain airfoils are given in figures 5 and 6, respectively. The data for the N. A.
C. A. 23012 and 23021 airfoils are discussed in references 1 and 3, respectively, and therefore require no further .
discussion. The data for the N. A. C° A. 23030 airfoil, I I I I I I I I I I_---- .012 however, depart from the results of the thinner sections _._, _ __, k in several respects. The slope of the lift curve is only 0.068 as compared with about 0.105 for the N. A. C. A.
23012, although there is a marked increase in slope at angles of attack above 2 ° • The angle of attack for zero lift, however, is the same as for the N. A. C. A. _n ,._z 28 32 P 23012 and 23021 airfoils. The relatively flat-top lift A/rfoil th/ckne_s, percent cjSZA C.A. 230,ser s) % curve given by the N. A. C. A. 23030 airfoil is probably typical of very tt_ick airfoils. Its pitctmlg-momeIlt Fl(_uR_7.--Etleetoftbicknessofplainairf°ils°nmaximumliltandminimumdrag" REPORT NO. 668_NATIONAL ADVISORY COMMI'FrEE FOR AERONAUTICS ]:"uultE 8.--Sectiou aerodynamic characteristics of N. &. C.._.. _012 airfoil with various sizes of split flap.
23030 AIRFOILS WITH SPLIT FLAPS N. A. C. A. 23012, 23021, AND I _.4 2.8 -/.2 -.8 -..4 0 .4 .8 1.2 /.6 0 .4 .8 Sect on/ift coefficient, cz (d) The 0.40c_ split flap.
(c) The 0.30c _ split flap. .
Ftouaz 8.--Continued. Seetiun aerodynamic characteristics of N. A. C. A. 23012 airfoil with various sizes o[ split flap.
REPORT NO. 668_NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 0 .4 .8 Z2 L6 20 2.4 2.8 */.Z -8 =4 0 .4 .8 LZ /.6 Sechon lift coefficienf, c_ (a) The 0.10c. split Uap. (b)The 0.20c. split flap.
FIGURE 9.--Section aerodynamic characteristics of N. A. C. A. 23021 airfoil with various sizes ofsplit flap.
N. A. C. A. 23012, 23021, AND 23030 AIRFOILS WITH SPLIT FLAPS | | • / -.4 0 .4 .8 /.2 1.6 2.0 2.4 8 -/2 -,8 -:4 ,Sect on tiff coeff/cienf, cz (c) 'I'ho 0.30cw split ilap. (d) The 0.40c_ split iiap.
Fic, uRz 9.--Continued. Sectionaerodynamic characteristics of N. A. C. A. 2.3021 airfoil with v_rious sizes ofsplit flRP.
16156_--39----2 REPORT NO. 668--NATIONAL &DVISORY COMMITTEE FOR AERONAUTICS • 4O 0 .4 .8 LE L6 2..0 2.4 2.8 -Z2 -.8 -_4 0 .4 .8 /.2 /.6 Seclion /iffcoefficieni, c t (a) The 0.10e. split flaps. (b) The 0.20e. Split flal)s.
FrciL'U_ lO.--Section aerodynamic characteristics of N.._...C.._.. _030 airfoil with various sizes of spilt flap _. ._.. C. A° 23012, 23021, AI_ID 23030 AIRFOILS WITH SPLIT FLAPS 4- ..n -- m • O4 (d) The 0.40_ split flap.
(c) The.0.30_,_ split flap.
b't_ug_ 10.--Continued. Section aerodynamic characteristics of N.._.. C. A. 23030 airfoil with various sizes o[ split itat_.
10 REPORT NO. 668--NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS Airfoils with flaps.--The section aerodynanfic _;har- The effects of a change in thickness of the plain air- acteristies of the N. A. C. A. 23012 airfoil with the foils on the minimum profile-drag coefficients and on the 0.10c_, the 0.20c_, the 0.30e_, and the 0.40cw split maximum lift coefficients are indicated in figure 7 for flaps are shown in figure 8. All these data were an effective Reynolds Number of 3,500,000. Although obtained at an effective Reynold_ Number of 3,500,000, the minimum profile-drag coefficient increases rapidly except as noted on the figure. The lift curves have with airfoil thickness and is nearly twice as great for about the same slopes as they did for the plain airfoils.
the N. A. C. A. 23030 as for the N. A. C. A. 23012 The angle of attack for maximum lift decreases from airfoil (see fig. 7), it may be that structural considera- /.2 1.6 2..0 2.4 2.8 _4 0 .4 .8 /2 /.6 20 2.4 2_ _cect/on I/ft coefficient, c t (a) The 0.10c, split flap. (b) The 0.20c, split flap.
(c) The 0.30¢,, split flap. (d) The 0.40e, split flap.
FI(;URg ll.--Comparison of profile-drag coefficients for airfoils with split flaps• about 15 ° with the flap neutral to about 14 ° with the tions will more than overbalance this drag increase in flap down 30 °. With the flap down 60 ° or 75 °, how- application to a given design. In other words, the ever, the angle of attack for maximum lift is only about probability should not be overlooked of actually 10 ° or 12 °, a change of 5 ° or 3 ° from the plain airfoil.
obtaining desired characteristics with the tlfick sections Changes of this magnitude in the angle of attack for because of the possibility of housing parts of tile maximum lift might have considerable effect on the airplane entirely within the wing, which would be manner in wllich a wing stalls for combinations With partial-span flaps.
iinpussible with the tt_mer sections.
N. A. C. A. 23012, 23021, AND 23030 AIRFOILS WITH SPLIT FLAPS ift Coefficients less than 1.8; for lift coefficients greater Similar section aerodynamic data are given for the than 1.8, it is lowest for the N. A. C. A. 23021 airfoil.
N. A. C. A. 23021 airfoil with flaps in figure 9 and for The drag is lowest for the N. A. C. A. 23012 airfoil the N. A. C. A. 23030 airfoil with flaps in figure 10.
with the 0.30c_ and the 0.40c_ split flaps for lift coeffi- The angle of attack for maximum lift with the thicker cients less than about 2.1; for lift coefficients greater airfoils with the flap deflected decreases with increasing than 2.1, it is lowest for the N. A. C. A. 23021 airfoil.
thickness and flap chord to values as low as 5°, a change With the 0.30c_ and the 0.40c_ flaps, the drag is lower of about 10 ° from the plain airfoil. It should also. be for the N. A. C. A. 23030 than for the N. A. C. A. 23012 noted that a considerable increase in the profile-drag airfoil for lift coefficients above 2.5.
coefficient is obtained with increase in the flap chord.
A comparison of the parts of figure 11 shows the The pitching-moment coefficient about the aerody- drag coefficients to be lowest for the smallest-chord namic center increases quite rapidly with flap chord, flap suitable for a given lift coefficient for take-off.
flap deflection, and airfoil thickness. The marked 40 60 80 0 80 40 60 80 0 20 40 60 80 /00 0 2O Flop deflect/on, '_.,,,, deg.
(a) N.._. O. A., 23012 airfoil. (b) N. A. O. A., 23021 airfoil. (c) N./_. C. A., 23030 airfoil.
FIGURE 12.--Effect of split-flap deflection on increment of maximum lift coefficient for the various airfoils and flaps.
All the combinations with the split flap have higher increase with airfoil thickness is probably caused by drag coefficients throughout the take-off range than the fact that the aerodynamic center is unusually far do the combinations with slotted flaps, which were above the chord line and ahead of the quarter-chord developed for the N. A. C. A. 23012 and 23021 airfoils point for the thick airfoils.
and are reported in references 1 and 3.
Effect on maximum lift._The effect of deflecting the COMPARISON OF AIRFOILS WITH FLAPS split flaps on the incremeut of sectiott maxinlunl lift Effect on profile drag.--The effect of the 0.10c_ split coefficient AC_,,a= is shown in figure 12, where hcl,,,,,_ is tiap ()it the profile drag of the three airfoils for various plotted against _t for all the combinations tested. The tlap deflections is shown as envelope polar curves in maximum Ac_,= increases with airfoil thickness for _tll figure 11 (a). Similar curves for the 0.20c_, tile 0.30c_, the flap chords. The flap deflection for maxinmm and the 0.40c_ flaps are given, respectively, in figures 5c_,= decreases with increase in flap chord for _my of 11 (b), 11 (c), and 11 (d). With the 0.10c_ flap, the the three airfoils. In figure 13, the ma_mum Ac_,, drag is lowest throughout the complete lift range for the N. A. C. A. 23012 airfoil. The drag is lowest for is plotted for each flap against flap chord for the three the N. A. C. A. 23012 airfoil with tim 0.20c_ flaps for
12 REPORT NO. 668JNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
for the airfoils with flaps. The final maximum lift airfoils. The highest Acz,_ for the N. C. A. C. 23012 coefficients for the N. A. O. A. 23012 and 23021 airfoils airfoil was obtained with the 0.30e_ flap, which was with the 0.10c_ flap was 2.34, which is about 8 percent only slightly superior to the 0.20e, flap on this airfoil.
higher than it was for the N. A. C. A. 23030 airfoil.
•The highest Ae_ for both the N. A. C. A. 2302Tand The maximum lift coefficient for the airfoils with the the N. A. C. A. 23030 airfoils was obtained with the 0.20c, flap was 2.66 for the N. A. C. A. 23021 airfoil, which is about 4 percent higher than it was for the I o " ' i 23030, N. A. C. A. 23012 and 2 percent higher than it was for the N. A. C° A. 23030 airfoil. For the airfoils with the /.6 0.30e_ and the 0.40e_ flaps, the maximum lift coeffieien_ was 2.6 for the 23012 airfoil and increased about 11 percent with airfoil thickness for the 21-percent-thick
/ 23021--
airfoil. The maximum lift decreased slightly with t-"_o thickness for the 0.30e_ flap and increased slightly for _a the 0.40c_ flap. The highest maximum lift coefficient, ._. 1.4 2.94, was obtained with the 0.40c_ flap on the N. A.
(J C. A. 23030 airfoil. In spite of the loss in lift of the 80* plain airfoils with thickness, if for structural reasons I 1.2 wing thicknesses are increased to as much as 30 percent, 95" i no loss in ultimate section maximum lift coefficient
////
_- Lo I
/ zs"
r _-- _ _ ' i -. 40c_ .9 _ _. 310C_ 80 ° J i_--..-- _....,..__.,.,_.._ f q_ _. _Oc_
/
h
ill
"" .lOcm _ .4
i
Plain oirfoi/ 0 I0 20 30 4O 50 F/ap chord, percent cw FlauRI 13.--Effect el chord of split flap on increment of maximum lift coefficient for three akfotl thicknesses.
0.40ew flap. The 0.40e. flap, however, gave little gain over the 0.30e, flap, and probably no gain would be obtained by the use of a flap chord greater than 0.40c_ on the N. A. (]. A. 23021 airfoil; for the No A. 0. A.
23030 airfoil, flaps of still larger chord might give a 12 16 20 24 28 32 slight increase in hcl,,,,,_.
Airfoil thickness, percent cw(N. A.C A. 230 metres) The increments _: maximum llft coefficient increase- FTOURE 14.--Effect of airfoil thlcknea_ on maximum lift coefficient of N. A. C. £.
quite markedly with airfoil thickness; the values of 230 airfoils with and without split flaps.
Ae_,_ vary from 1.05 for the N. A. C. A. 23012 to 1.9 will be encountered when split flaps with chords of for the N. A. C. A. 23030 airfoil. The final maximum 0.20e_ or larger are used.
lift coefficient, however, does not reflect this large differ- ence in Ae_a_, as is shown in figure 14, where el_a_ for SCALE EFFECT the plain airfoils and for the airfoils with flaps is plotted • The scale effect on maximum lift coefficients for the against airfoil thickness. The large loss in lift with plain airfoils and the airfoils with flaps, over the range thickness for the plain airfoil very nearly balances the available in the 7- by 10-foot wind tunnel, is shown in large gain in increment of maximum lift with thickness N. A. C. A. 23012, 23021, AND 23030 AIRFOILS WITH SPLIT FLAPS flapwere about equal; for the airfoils with the 0.10c_ figure 15, where ctm. , is plotted against the value of R, flap, the maximum lift coefficient decreased with airfoil thickness; and for the airfoils with the 0.30c_ and the of the tests. This figure shows a very definite scale 0.40e_ flaps, the maximum lift coefficient increased with effect on the maximum lift coefficient for the "N. A.
C. A. 23012 airfoil with or without flaps but shows airfoil thickness.
Within the range covered, the increment of maximum practically none for the N. A. C. A. 23021 and 23030 lift coefficient due to the split flaps was practically airfoils with or without flaps. The increment of maxi- mum lift coefficient is therefore practically independent independent of scale. The profile-drag coefficient in- of scale over the range that could be investigated. creased quite rapidly with thickness for the plain airfoils and was about twice' as large for the N. A. O. A.
APPLICATION OF OTHER AIRFOILS 23030 asfor the 23012 airfoil.
The maximum lift coefficients for airfoils of the N. A. C. A. 430 and 630 series with split flaps may be computed with satisfactory accuracy by adding the LANGLEY MEMORIAL .-_k-ERONAUTICAL LABORATORY, NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS, LANGLEY FIELD, VA., March 10, 1939.
, 0.40c,o flops REFERENCES 1. Wenzinger, Carl J., and Harris, Thomas A.: Wind-Tunnel Investigation of an N. A. C. A. 23012 Airfoil with Various .u Arrangements of Slotted Flaps. T. R. No. 664, N. A.
C. A., 1939.
_/.,5 _j 2. Harris, Thomas A.: The 7 by 10 Foot Wfnd Tunnel of the National Advisory Committee for Aeronautics. T. R.
No. 412, N. A. C. A., 1931.
3. Wenzinger, Carl J., and Harris, Thomas A.: Wind-Tunnel Investigation of an N. A. C. A. 23021 Airfoil with Various Arrangements of Slotted Flaps. T. R. No. 677, N. A.
C. A., 1939.
4. Jaeebs, Eastman N., Pinkertmnp Robert M., and Greenberg, Harry: Tests of Related Forward-Camber Airfoils in the Variable-Density Wind Tunnel. T. R. No. 610, N. A.
/ t/lll.liiiitlllll IIIIII
C. A., 1937.
i Z 3 4 5 ,, 8xlt TABLE I Effecfive Reymolds Number, fLe ORDINATES FOR N. A. C. A. 230 AIRFOILS FIG_E 15.--Effect of scale on maximum lift coemcicnt of three akfofls with and without split flaPS; 7- by 10-foot wind tunnel.
[Stations and ordinate6 in percent of wing chord] 23030 23021 Ac_ma, for the proper flap chord and airfoil thickness 23012 from the 230 series to the czma, of the plain airfoil under [3pper I Lower Station Upper I Lower 3"PPer I Lower _urfaCe surface ;urface surface urface I surface consideration. This procedure is justified for thick- 4.82 o nesses from 9 to 21 percent, as indicated in reference 4.
o 7.37 --2.63 "'i'_'" -2.os "_'_"1 -1.= 8.90 --4.27 The same procedure would also probably be satisfactory S. 14 --3.14 3.61 _ --1.71 11.05 --6.54 7.93 --4.52 4.91 --2.26 for other airfoils with the position of maximum camber 12. 57 --8. 28 9.13 --5.55 5. 80 --2. 01 13. 68 --9.65 10. 03 --6.32 6.43 --2.92 near the leading edge. It should be remembered in 15. 20 --Ii. 52 II. 19 --7. 51 7.19 --3. 50 16.07 --12.61 11.80 --8. 30 7. 50 --3. 97 16.46 --13.20 applying these data that they are section character- 12.05 --8.76 7.60 --4.28 16.57 --13. 46 12. 06 --8. 95 7.55 --4.46 istics and that these maximum lift coefficients cannot 15.89 --13.13 11. 49 --8. 83 7.14 --4.48 14.38 --12. Ii 10. 40 --8.14 be realized on a wing of finite span unless it is designed 30:--:-::--::::::- 0. 41 --4.17 12. 34 --i0.47 8.90 --7.07 5.47 --3.67 9. o_ --8. 42 7.09 --5.72 4.36 --3.00 7.03 --6.09 5.05 --4. 13 3. 08 --2.10 3.87 --3.40 2. 76 --2. 30 1.68 --1.23 2.15 --1.80 1.53 --1.30 CONCLUDING REMARKS _ _----_-_-------_--'--: .92 --.70 .32 --. 32 .22 --.22
.I__L l
9.90 4.85 1.58 Aerodynamic data are made available for airfoils 12 = ................. E. radius ........
so that all sections reach maximum lift s;.multaneously, i/::::::::::::::::::::: to 30 percent thick with split flaps having chords 10 to Slope of radius through end of clmrd: 0.305 40 percent of the wing chord. The final maximum lift - coefficients for the three airfoils tested with the 0.20c_ U.$.GOVERNMENT pRINTING OFFiCE:t939
j
Z Positive directions of axes and angles (forces and moments) are shown by arrows Velocities Axis Moment about axis Angle Force Linear (parallel (compo- Designa- Sym- Sym- to axis) Designation Sym- Positive Angular Designation nent along tion bol bol symbol bol direction axis) I Rolling____ _ L Y----> Z Roll_ _ _ __ tt Longitudinal ..... X X P Lateral ........ : _ Pitching_:__, M_ Z---_X q r Normal .......... Z Z Yaw ..... ¢ Yawing_._ 2_ N X---_ Y Absolute coefficients of moment Angle of set of control surface (relative to neutral position), _. (Indicate surface by proper subscript.)
C_q_S M --_lbS (rolling) (pitching) (yawing) 4. PROPELLER SYMBOLS P Diameter , P, Power, absolute coefficient Cp_p_-D6 Geometric pitch P, • 5pF _ Pitch ratio p/D, C_, Speed,power coefficmnt =,_/F_ V', Inflow velocity 7, Efficiency
V,, Slipstream velocity
n, Revolutions per second, r.p.s, G-- T r, Thrust, absolute coefficient T--pn-VD a Effective helix angle=tan-l(_ v _ \z_rn/ Q_ Torque, absolute coefficient Cq----p---_r_,._ _. NUMERICAL RELATIONS 1 hp.--76.04 kg-m/s----550 ft-lb./sec. 1 lb.=0.4536 kg.
1 metric horsepower=l.0132 hp. 1 kg----2.2046 lb.
1 m.p.h.=0.4470 m.p.s. 1 mi.--1,609.35 m----5,280 ft.
1 m----3.2808 ft.
1 m.p.s.=2.2369 m.p.h.