Document
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MLMORANDUM
RESEARCH
' for the Air Materiel Command, Army Air Forces THE EFFECTIVENESS AT HIGH SPEEDS OF A 20—PERCENT—CHORD PLAIN TRAILING—EDGE FLA_ p ON THE NACA 65-210 AIRFOIL SECTION By Louis S. Stivers, Jr.
Ames Aeronautical Laboratory Moffett Field, Calif.
Restriction/Classification Cancelled This d e National Defense of the United States within the meaning of the Espionage Act, USC 50:31 and 32. Its transmission or tha revelation of its contents in any manner to an unauthorized person is prohibited by law. Information so c!assified m.,.y be imparted only to persons in the mi!itary and naval Services of the United State appropriate civilian officers and employees of the Federal Governmc.r.t who have a legitimate inter,ist therein, and to United States citizens of known Icyaity <4.id scretion who of necessity rrr , - mform 7 ^] r.."
NATIONAL ADVISORY COMMITTEE
FOR AERONAUTICS
WASHINGTON
MAY 6
19-".7
t
13;. NI T I it W' - l NACA RM No. A71,17 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS RESEARCH MCMORANDUM for the Air Materiel Command, Army Air Forces HE EFFECTIVETTF .ESS AT HIGH SPEEDS OF A 20 PERCETTT-CHORD -210 AIRFOIL PLAID? TRAILING-EDGE FLAP DDT THE , DTACA 65 SECTION By Louis S. Stivers, Jr.
SUMMARY An analysis ha-s been made of the lift-control effectiveness of a. 20-percent-chord plain trailing-edge flap on the NACA 66-210 airfoil section from section lift-coefficient data obtained at 0.875. In addition, the effectiveness Mach numbers from 0.3 to of the plain flap as a. lift-control device has been compared with the corresponding effectiveness of both a spoiler and a dive- recovery flap on the DTACA 65-210 airfoil section.
The analysis indicates that the .plain trailing-edge flap employed on the 10-percent-thick airfoil at Mach numbers as high as 0.875 retains at least 50 percent of its low-speed lift-control effectiveness, end is sufficiently effective in late2a1 control application, assuming a rigid wing, to provide adequate airplane rolling characteristics.
The plain trailing-edge flap, as compared to the spoiler and the dive-recovery flap, appears to afford the most favorable characteristics as a device for controlling lift continuously 0.875.
throughout the range of Mach numbers from 0.3 to At Mach numbers above those for lift divergence of the wing, either a plain flap or a. dive-recovery flap may be used on a thin airplane wing to provide auxiliary wing lift when the airplane is to be controlled in flight, other than in dives, at these Mach numbers. The choice of a lift-control device for this use, however, should include the consideration of other factors such as the J 2 _. NACA RM No. A7A17 increments of drag and pitching moment accompanying the use of the device, and the structural and high-speed aerodynamic character- istics of the airplane which is to employ the device. - INTRODUCTION Among many effects of compressibility which have been found in flight and in the wind tunnel is a large reduction in the effectiveness of conventional airplane control surfaces at veloc- ities considerably above the airfoil critical speeds. In some instances the effectiveness has been shown to reduce to nearly zero at nigh speeds, thus definitely limiting the maximum speed of controlled flight. In order to determine whether this reduction in effectiveness is influenced by the type of control surface employed, various lift-control devices on relatively thin airfoils have been investigated at high speeds.
The lift-control effectiveness of spoilers and dive-recovery flaps used on thin airfoils has been reported in references 1 and 2.
The spoilers became decreasingly effective with increasing projection at high Mach numbers, and exhibited characteristics which were such as to promote erratic lift control at high speeds. The dive- recovery flaps also showed generally unfavorable characteristics for use, other than emergency, as lift control devices at high speeds.
Wind-tunnel data, presented in reference 3 for a. plain trailing -©dge series airfoil 19 percent thick indicated flap on a. modified TTACA 6- that the effectiveness of a. plain flap used for lateral control on a thick airfoil rapidly decreases as the Mach number is increased above the a.irf oil critical Mach number.
In order to provide information on the lift-control effective- ness of a plain trailing-edge flap on a, representative thin PTACA 6- series airfoil, the present analysis was undertaken. For comparative purposes,increments of section lift, dra,.g,and pitching- moment coefficients for the plain flap together with the correspond- ing characteristics of the spoiler and the dive-recovery flap are presented for the range of Mach numbers from 0.3 to 0.877. The analysis pertaining to the plain flap was made using data from reference 4. the effect of the differences in rigidity of the wind-tunnel models and the various parts of an airplane has net been considered in the present analysis.
NACr. RM No.
A7117 • SD,[BOLS cZ section lift coefficient ' nscz increment or decrement in section lift coefficient Lca increment in section drag coefficient
6 N .cm C/4 increment in section moment coefficiont about quarter—
chord point M free—stream Mach n.umber Me section angle of attack, deE;°ees flap deflection, degrees
b f
' ^'Qo/A3f section lateral—control—effectiveness parameter, absolute value of the ratio of equivalent change in section angle of attack to change in flap deflection angle at a constant section lift coeffic-.ent METHODS OF ADLILYS IS The present analysis of flap effectiveness was made using a.erodyr_amic data obtained in the lines 1— by 3 foot high.—speed wind tunnel from tests of the NAC_^ 65-210 airfoil equipped with a.
20—percent—chora plain flap. These data were obtained for speeds ranging from 0.3 to approximately 0.9 Mach number (with a. corre-
sponding range in Reynolds numbers from a.pproxima.tely 1 x 10 6 to
2 x 10 6 2 0 to 80 and flap
) for airfoil angles of attack from — More precisely, the flap deflections deflections from — 6 0 to 6 0 .
2.6, 0, 1.9, 4.6, and 6.3.
in degrees were found to be —6.3, —4.9, —
The lift—coefficient data. for a Mach number of approximately 0.9
were not obtained at a. sufficient number of airfoil angles of
attack to permit their use in the present analysis. For this reason, only data, for Mach numbers as high as a..ppea.r in the figvxes.
0.8'75 In order to indicate tho effectiveness of the plain flap as a.
lift producing device, increments of section lift coefficient for
each angle of flap deflection have been determined. These increments
were obtained throughout the Mach number range at airfo i l angles of
attack corresponding to lift coefficients of 0, 0.2, 0.4, 0.6 ,and 0.8
at zero flap deflection. Fcired curves showing these increments for
constant Mach n7uanbers are presented in figure 1 as a, function of flap
4 I _ NIX"' R4 No . A71,17 deflection. The sane increments for constant flap deflection cross- plotted at cash rf a.i oil angle of attack given in figure 1 are presented in figure 2 as a. function of Mach number.
The effectiveness of a lateral-control device is not ir_dica.ted
completely by increments of lift coefficient alone. Some parameter must be used which considers the changes in airfoil lift-curve slope with changes in control surf ace deflection. The commonly used pa.rn: meter '!^ao/Abf, defined as the ratio of the change in airfoil-section.
angle of n.tta.ck to the char-go in flap deflection necessary to main-
tain a. constant lift coefficient, has been adopted for use in tho present analysis. The variation of this parameter with Mach number for the plain flap of the present report is given in figure 3 for several moderate lift coefficients. For comparison, the variation
of the lateral--control-effectiveness parameter with Mach number
for a. 20 percent--chord plain fla.p on a 19-percent-thick modified.
NIXA 6i-series airfoil is also shown in figure 3. The curve for
4+ the latter airfoil and flap was obtained from figure of reference 3.
For the present report, values of Aao/GSf were taken as the absolute
value of the average slopes of the curve of section angle of attack
versus flap deflection over a. range of flap deflections from --6 o to
6 0 , for a. constant section lift coefficient.
A graph (fig. 4) has been prepared which illustrates the respective variations with Mach number of increments in. section lift coefficient with flap deflection for the plain flap and for the dive-recovery flap, and of decrements in section lift coeffi- a. spoiler. From the hiE41-speed investiga- cient with projection for tion (two-dimensional) of a spoiler loc-ated at several positions on the upper surface of the NACA 65-210 airfoil section, it appeared that the 50-percent-chord location was the most suitable investigated.
Decrements of lift coefficient for various spoiler projections at
this location are shown in figure 4 for an airfoil angle of attack
corresponding to a. lift coefficient of 0.2 at zero spoiler projection.
Similarl 7 , the increments of lift coefficient for several (live-
recovery flap deflections are also shown in figure 4 for a correspond-
ing airfoil angle of attack and for the dive-recovery flap located
at the 50-percent-chord position. The high-r^peed investigation (two-
dimensional) of dive-•recovery flaps indicated that, of three flap
locations on the lower surface of the MTA 65-210 airfoil, the 50-percent-chord position wa.s also the most suitable location.
The changes in section drag a.nd. pitching-moment coefficients corrosponding to the increments (or decrements) of lift coeff i--
ciant shown in figure 4 are presented in figures 5 and 6, respec-
tively, for the same three lift-control devices.
NACA . RM No . A71,17 . l 5 The dotted portions of certain curves appearing in figures 1 and 2, and of the curve of figure 3 for the 19— percent—thick airfoil are used to indicate that some uncertainty exists regardinb the validity of those data obtained in the vicinity of the wind—tunnel (0.9 choking Mach number at zero angle of attack for the NACI.
0.74 at zero angle of attack 65-210 airfoil model, and approximately for the 19-percent—thick airfoil model) .
DISCUSSION h desirable lift—control device for use on aircraft wings or tail surfaces is one which has uniform effectiveness throughout the range of Mach nuribers at which the device is expected to be employed. Fu rthermore, if an airplane is to maintain controlled- flight at Mash nmbers above those for lift divergence of the wing (which axe generally lower than those for lift divergence of the tail), it must be possible to compensate for the lift deficiency of the wing at these Mach numbers. These two particulars are considered in the succeeding discussion both in regard to the plain fla.p of the present analysis and in regard to the comparison that follows. The two—diriensicna.l data. presented herein can indicate, in general, the aerodynamic effects on on airplane wins; or tail resulting from the use of one of the lift—control devices.
which It should be remembered, however, that several other factors are not considered in this analysis, such -is the difforences in the aerodynamic cha.racteristics of the tail a.nd. wing, the downwa.sh at the tail, and the elevator hinge—moment chara.cteristics, may greatly affect the over—all longitudinal—stability a.nd ---control characteristics of an airplane in flight, especially at high speeds.
Effectiveness of the Plain Flap as a. Lift—Producing Device The increments of section lift coefficient shown in figures 1 and 2, which indicate the effectiveness of the flap a.s a. lift— producing device, show that the effectiveness increases somewhat with increase in Mach number reaching a. maximum at a. Mach nlamber apparently depending on the ma.gnitiide of the flap deflection and the airfoil angle of attack. The Mach numbers for which the increments of lift coefficient are greatest correspond approxi- mately, in must cases, to the airfoil lift—divergence Mach numbers 4, In the range of 14a.ch numbers given in figure 8 of reference from those at whic?m the maximum increments occur to 0.875 Mach rnarmber the effectiveness decreases in varying degree, The min in. url effectiveness indicated, however, is never less than 50 percent of NACA RM No. A7A17 that at low speeds. Although the data of figures 1 and 2 indicate appreciable variations in the effectiveness of the plain flap for Mach numbers between 0.3 and 0.875, it is believed that these variations will not too seriously limit the application of this control device on a. 10-percent-a-iick rigid airfoil in the said Mach number range.
Figures 1 and 2 further indicate the plain flap to be capable
of providing substantial increments of lift coefficient for small
flap deflections at Mach numbers above those for airfoil lift
divergence. The plain flap, then, used either on a. thin rigid air-
plane wing or tail -remains effective as a. lift-producing device at
speeds greater than those corresponding to the wing or tail lift divergence, respectively.
Effectiveness of the Plain Flap for Lateral Control
The lift-control characteristics of a plain flap at high, speeds
are of further significance from the standpoint of the lateral
control of an airplane. The lateral-control effectiveness of the
plain flap of the present report can be evaluated from the data of
figure 3 which show the variation_ with Mach number of the section
ia.tera.l-cor_t-rol-effectiveness parameter c^a.o%af. . For any given
airplane the magnitude of the parameter pb/2V (helix angle generated by the wing tip of an airplane in roll) is directly proportional to the airfoil-section lateral--control parameter
Aa, (assuming a. rigid airplane wing). A study of the variations
ob N3 f
Of ZyMo/.!\sf with Mach number will, accordingly, correspond to a.
study of the variations of pb/2V of an airplane employing the air- foil and lift-control device. Furthermore, whatever decrease in the values of pb/2V with increaso in Mach number can be allowed for an airplane, consistent with the maintenance of adequate lateral control, can also be allowed for the a.irfoil-section parameter ^a.o lAb f .
The data. of f i.gizre 3 for the NACA 65-210 airfoil with a. plain
flap show an appreciable variation in lateral-control effectiveness over a. range of moderate lift coefficients at nigh Mach numbers.
The only marked decreases in effectiveness, however, appear to
begin at Mach numbers near 0.83 for low lift coefficients. he
largest decrease in effectivenoss, for Mach numbers up to 0.875, is indicated for zero lift coefficient where the effectiveness has reduced to a. value which is approximately 50 percent of that shown for low speeds. In a. Navy Department specification for the stability a.nd control characteristics of airplanes (reference 5), C NACA RM No. A7A17 no reduction in the minimum allowable va.lue of pb/2V for adequate lateral control is permitted for indicated airspeeds up to 300 miles per hour, but a two-thirds reduction is permitted for an increase in indicated airspeed from 300 to 500 miles per hour. At an 5 at altitude of 10,000 feet (an altitude specified in reference which compliance with these lateral-control requirements are to be demonstrated bzr the airplane in flig_7t) indicated airspeeds of 300 and 500 miles per hour correspond, respectively, to approximately 0.5 and 0.8 Mach numbers. The plain trailing-edge flap applied to a rigid wing appears, then, to exhibit adequate lateral-control character- istics up to Mach numbers as high as 0.875.
for the two airfoils A comparison of the curves of figure 3 employing 20-percent--chord flaps shows that the effectiveness exhibited by the flap on the 19-percent-thick airfoil at high speeds is quite different from that fur the 10-percent-thick airfoil.
The curve for the 19 percent-thick airfoil shows a. marked decrease in the effectiveness of the flap at a. Mach number near 0.70 which is approximately 0.13 Mach number less than that corresponding to the abrupt decrease in effectiveness of the flap on the 10-percent- thick airfoil at low lift coefficients. It can also be noted from the data of figure 3 that, while serious losses in the effectiveness of a. flap on a. 19-percent-thick airfoil ca.n be expected above Mach numbers of th-- order of 0.7, no severe losses should be expected for a, plain flan on a. 10 percent-thick a.irfoil, especially for higher lift coefficients, up to Mach numbers approaching 0.875.
Comparison of the Lift-Control Effectiveness of a Spoiler, Divo Rocovery Flap, and a. Plain Flap The relative merits of a spoiler, a. dive-recovery flap, and a.
plain flap for providing lift control on an a.irf oil can be evaluated 4. It can be from the lift coefficient data. presented in figure with Mach number of seen re-adily from the data that the variations the lift-control effectiveness of the spoiler and the dive-recovery to 0.875 are considerably larger flap from a. Mach number of 0,3 than the corresponding variations for the plain flap, Because of these large variations in effectiveness for the dive-recovery flap, a.nd especially for the spoiler, an airplane control system employ- ing either of these devices would tend to provide at high speeds too rapid airplane response to control movements if sa.tisfa.ctory low- speed control characteristics were maintained. For producing lift continuously throughout a. wide range of Mach numbers, the plain trailing-edge flap, accordingly, appears to possess the most fa.vo:°- able chara.cteristics.
8 NACA RH No. A7A17 For providing auxiliary lift at Mach numbers above those for airfoil lift divergence, the plain flap deflected in a, positive sense and the dive—recovery flap are considered for positive increments of lift; whereas the plain flap deflected in a negative sense and the spoiler, on the other hand, are considered for negative incre- ments of lift. The data. of figure 4 show that each of these lift devices is capable of providing increments (or decrements) of lift 0.875.
coefficient in the range of Mach numbers between_ 0.77 and (This range includes Mach numbers above those for lift divergence of the airfoil). These increments, however, vary differently for each lift device with changes in Ma.cl: number and decrease with increase in Mach number a.t the highest Mach numbers shown, except for the 100 deflection of the dive—recovery flap and for positive deflections of the plain flap. The plain flap appears to have no particular advantage over the dive—recovery flap for providing positive incre- 0.75 and 0.875 on a 10—percent- ments of lift at Mach numbers between thick airfoil unless it be at the highest Mach numbers. At the Mach numbers near 0.875 the data for the plain flap show that the increments of lift coefficient for the larger flap deflections do not continue to decrease with increase in Mach number as the corresponding incre- ments do for the dive—recovery flap.
The increments of drag coefficient corresponding to constant increments of lift coefficient, as shown in figure 5, are seen to be quite different for the three lift—control devices. The characteristics for the plain flap appear to be the most desirable, since the data indicate that the increments in drag accompanying a. given increment in lift is the lca.st for the plain flap at any Mach number from 0.3 to 0.875. Between 0.75 and 0.875 Mach numbers the increments in drag coefficient for constant increments of lift coefficient of the dive— recovery flap increase very rapidly with increase in Mach number.
In the case whore a. lift—control device is used on an airplane wing as purely emergency implement for aid in recovery from high—speed dives, a substantial increase in dra.g, such as noted for the dive— recovery flap, may be desirable in order to limit the diving speed of the airplane.
At constant increments of lift coefficient, the i ncrements of pitching—moment coefficient presented in figure 6 do net vary a great deal with change in Mach number except, for the most part, at the highest Mach numbers. For negative increments of 7_ift a.t Mach numbers botween 0.3 and 0.875, the plain flap and the spoiler exhibit, in general, positive increments of pitching moment which tend to increase at the highest Mach numbers for the larger negative incre- ments of lift. The pitching moment increments for positive incre- ments of lift are negative for the plain. flap, and positive for the NACA RM No. A7A17 4 9 dive—recovery flap except for the larger increments of lift at high increments of pitching moment Mach numbers. The data, show that the
are always more positive for the dive-recovery flap than are the
corresponding pitching--moment increments for t17o plain flap. In
the range of 14ach numbers from 0.75 to 0.875 the pitching--3^oment
coefficients for the plain flap are always negative (not in the
diroction to oppose the diving tendency); whereas for the di.vo-
recovery flap they appear to be either positive or negative, depend- ing on the Mach number and the increment of lift coefficient. A
negative increment of pitching mor.iient accompanying the use of any
control device at high subsonic speeds should certainly be consid- lift - ered in the structural and- aerodynamic design_ of an airplane tail.
T'he diving tendency of airplanes, resulting from the loss in wing lift at Mach numbers above those for lift divergence of the wing, lity an increase in longitudinal stabi
is generally accompanied b y
: changes. As a consequence, the control forces of some
and by tri p
a.irpla.nes in high—speed dives increase to such an extent that it
necessary to employ dive-recovery flaps as an
has been found lot in pulling out from the dives.
emergency device to aid the p i
On the other hand, pilots of some of the _,iore recent high-speed aircraft have effected -recovery from high-speed dives without recourse to emergency devices. In emergency applications the dive-recovery
flaps are advantageous in that they increase the wing lift for air-
an increment of lift together with a favor-
plane trim by providing
t
or the dive-
able pull—out momen . Since the data of figure 6 f
recovery flap show that the pitching-moment increment is not always positive, it would appear that the use of theso flaps on an airplane wing may not always provide favorable pitching moments for dire recovery as the Mach number or increment of lift is increwod. (The data. of reference 2 show that the dive-recovery flap located on the airfoil as far forward as the 30-percent-chord position also provides negative increments of pitching moment at high subsonic Mach numbers, except when the flap has a. small chord ratio.)
If an airplane is to be controlled in flight, other than in dives, at P.ach nrr:lbers above those for lift divergence of the wing, the use of dive recovery _Fla.ps at these Mach numbers to provide auxiliary lift on the wind may be limited by the large increase in drag. The choice of a lift-control device for such operation should also depend upon a.. consideration of other factors such as the incre- ments of pitching moment accompanying th, use of tho device, end the structural and high-speed aerodynamic characteristics of the airplane which is to e:aploy the device, NACA RM No. A7A17 CONCLUSIONS Tile analysis of the lift—control characteristics of a 20—percent— chord plain trailing—edge flap on the DTACA 65-210 airfoil section and a. comparison of the effectiveness of this device with that of both the spoiler and the dive—recovery flap indicate the following: 1. At Mach numbers as high as 0.875, the plain flap on the 10—percent—thick airfoil retains at least 50 percent of its lo„- speed lift—control effectiveness, and is sufficiently effective in lateral control,assuming a rigid wing, to provide adequate air- plane -rolling characteristics.
2. As compared to the spoiler and the dive—recovery flap, the plain trailing—edge flap would appear to afford the most favorable characteristics as a device for controlling lift continuously 0.875.
throughout a range of Mach numbers from 0.3 to 3. An airplane employing thin wings which is to be controlled in flight, other than in dives, at Mach numbers above those for lift divergence of the wing may use either a plain_ flap or a dive—recovery flap at these Mach numbers to provide auxiliary lift on the wing. It should be -remembered, however, that the choice of a. device for this use should include the consideration of other factors such as the increments of drag and pitching moment accompanying the use of the device, and the structural and high---speed aerodynamic characteristics of the airplane for w_zich the choice is to be made.
Ames Aeronautical Laboratory, National Advisory Committee for Aeronautics, Moffett Field, Calif.
00 b ^ 11 NACA RM Wo . A7A17 REF=CES Velasco, Carlos E.: High—Speed Wind-Funnel. InvestiFa.tion of 1.
Spoilers for Lateral Control on the NACA 651-210 Airfoil Section. NACA CT T IM No. A5K02, 1945.
Olson, Robert N., ar-d Bonozra, Joseph N.: High—Speed Wind— 2.
Tunnel Investigation of Dive Recovery Flaps for Lift Control on the NACA 651-210 Airfoil. NACA ARR 6D23, 1 946, 3. Lindsey, W.F.: Effect of Compressibility on the Pressures and Airfoil Having a Forces Acting on a Modified NACA 65,1-0 19 0.20--Chord- Flap. NACA ACR No. L5G31a, 1946.
4. Graham, Donald J., and Adams, Charles N.: Wind—Tunnel Investiga- tion of a 20—Percent—Chord Plain ^ la.p on the NACA 651-210 Airfoil for Lift Control at High Speeds. NACA CPR No. A5F05, 1945.
5. Anon.i Specification for Stability and Control Characteristics of Airplanes. Spec. No. 119A, Bur. Aero., Navy Dept., Apr. 17, 1945•
i
_.^ DT_ACA RM No. A7A17 FIGURE LEGEIMS Figure 1.- Variation of the increment of section lift coefficient with flap deflection at various Mach numbers for several angles of attack of the NACA 65--210 airfoil with a. 0.20-chord flap.
Figure 1.- Concluded. NACA 65-210 airfoil with a. 0.20-chord plain flap.
Figure 2.- Variation of the increment cf section lift coefficient with Mach number for various flap deflections and angles of attack of the NACA 65-210 airfoil with a. 0.20-chord plain flap.
Figure 2.- Concluded. NACA 65-210 airfoil with a 0.20-chord plain flap.
Figure 3.- Comparison of the lateral-control effectiveness at various Mach numbers for the NACA 65-210 and 19 percent thick 65-series airfoils with 20-percent-chord plain flaps.
Figure 4.- Comparison of the lift-control characteristics of a.
spoiler, a dive-recovery flap, and a, plain flap on the NACA 65-210 airfoil section at an angle of attack corresponding to a lift coefficient of 0.2 for zero deflection of the control device„ Figure 5.- Comparison of the increments of section drag coefficient corresponding to constant values of increment in lift coefficient given by a. spoiler, a. dive--recovery flap, and a plain flap on the NACA 65-210 airfoil section at an angle of attack corresponding to a, lift coefficient of 0.2 for zero deflection of the control device.
Figure 6.- Comparison of the increments of section moment coeffi- cient corresponding to constant values of increment in lift coefficient given by a spoiler, a. dive-recovery flap, and a. plain flap on the NACA 65-210 airfoil section at an 3.ngle of attack corresponding to a. lift coefficient of 0.2 for zero deflection of the control device.
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