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RM No. A7F09

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RESEARCH MEMORANDUM

A SUMMARY AND ANALYSIS OF DATA ON DIVE-RECOVERY FLAPS

By Lee E. Boddy and Walter C. Williams

Ames Aeronautical Laboratory Moffett Field, Calif.

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NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

AVASHIN6TON

September 9, 1947 : N AC A LIBRAE

J^WGLEY MEMORIAL AF.30NAUTCCAI ^ LABORATORY

Lamdey Firirt. V».

Ill NASA Technical Ub«y

EM Wo. A7F09

NATIONAL ADVISORY COMMITTEE FOE AERONAUTICS

RESEARCH MEMORANDUM

A SUMMARY AND ANALYSIS OP DATA ON DiHE-RECOVERY FLAPS

By Lee E. Boddy and Walter C. Williams

SUMMARY

The results of numerous unrelated teste of dive—recovery flaps are collected in this report and presented in a form suitable for use in the preliminary design of dive—recovery flap installations. Since the data were obtained for airplane models of quite widely varying configurations, and are limited largely to a Mach number of 0.80, it is recommended that each new installation be carefully flight—tested before final approval. A flight-test procedure is outlined which will insure a maximum degree of safety.

INTRODUCTION

Considerable difficulty has been experienced with many modern conventional airplanes in recovering from high—speed dives. As a result, various corrective devices have been investigated, the most successful of which has been the dive—recovery flap. This device is a small split flap mounted on the lower surface of the airplane wing. A typical installation on a wind—tunnel model is shown in figure 1, and an experimental installation on an airplane for flight tests is shown in figure 2.

Dive—recovery flaps were tested first in the Ames l6-foot high- speed wind tunnel in October 19^2, and were first tested in flight by the Lockheed Aircraft Corporation from December 19k2 to April 19^3- Subsequent flight tests were made by tfae Army Air Forces and the Republic Aviation Corporation with a Republic P—kf airplane and by the Langley Memorial Aeronautical Laboratory with a North American XP—51 airplane. Mere recently the Ames l6~foot high-speed wind tunnel has tested dive-recovery flaps on a number of airplane models in conjunction with more general investigations.

It is the purpose of this report to collect all the available

RESTRICTED

NACA EM No. A7P09

data on the subject and present them in a form which will serve as a guide for the preliminary design of dive—recovery flaps. By necessity, many factors pertaining largely to airplane configuration have been ignored. Consequently, a flight—test procedure is recommended. It is believed that the data, presented herein, if used in conjunction with the recommended flight-test procedure, will facili- tate the developement of satisfactory dive—recovery-flap installations for most conventional airplanes.

SYMBOTS

The following symbols are used in this report:

General

V free—stream velocity, feet per second

p free—stream mass density, slugs per cubic foot

q free—stream dynamic pressure (feY2), pounds per square foot

M Mach number I — ] Velocity of sound./

£M increase of Mach number over that for lift divergence

P pressure coefficient

(local static pressure)—(free-stream static pressure) | L q J

Pcr critical pressure coefficient (P at which the local velocity equals the local velocity of sound)

g acceleration of gravity, foot per second per second

£E decrease of total pressure from freo—stream total pressure, pounds por Bquaro foot

Airplane or Model Dimensions

S wing area, square foet

FACA EM No. A7F09 3

Sf total flap area, square foot

"fa wing span, feot

"bf total flap span, feet

"bo elevator span., feet

M.A.C. vlng noan aerodynamic chord, feet

cv average wing chord at the flap, feot

cf average flap chord, feot

c02 noon—square olovator chord, square feet

x longitudinal distance from the wing leading odge, foot

Coefficients

CL lift coefficient ( H££ }

Cn pitching-moment coefficiont (pitching noBont^ ^ V qS M.A.C. J

^Cm increase of pitching-monent coefficient for constant lift coefficient due to the flaps

/Cm«- incroase of ving pitching-incoent coefficient for constant lift coefficiont duo to the flaps

££%) increase of drag coefficient due to the flaps /increase of drag \ V IS J

ar hinge racanont \ q CQ2 hQ /

Che elevator hinge-aooent coefficient /o^P^tor hjngo raoaont }

Angles

a airplane or model angle of attack, degrees

c&u. uncorrected angle of attack, degroes

Act increase of angle of attack for constant lift coefficient due to the flaps, degrees

NACA EM No. A7F09

at airplane or model tail anglo of attack, degrees

AaCffc increase of tail anglo of attack at constant lift coefficient duo to tho flaps3 degroea

6f flap deflection, degreos

Bs olovator dofloction, dogroos

&50Q incroaso of olovator floating anglo duo to tho flaps, dogroos

RESULTS

Sourco of Data

The wind—tunnol data colloctod in this report arc the results of numerous unrelated toste conducted in tho Amos 16-foot high-speed wind tunnol from October 19^2 to Juno 19^5• Sovcral improvements of the model support system wore iaada during this period, particularly with regard to increasing its critical Mach number and reducing its interference at high Mach numbors. Although all tho data have been recently corrected for taros, constriction., and flow inclination due to tho support system according to latest knowlodgo, seme discrepancies may bo present bocauso of tho varying intorferonco of different support systems and different flap positions rolativo to tho supports.

Tho flight data ero the results of tests conducted by tho Langloy Laboratory.

Presentation of Results

ttsa.— Tho offoct of dive—recovery flaps on the lift and pitching-moment characteristics of the models tested in tho wind tunnol is shown in figuros 3 to 1^. Included in oach figuro is a half plan viow of tho model as well as pertinent geometrical information on tho flap installation. Additional information concerning tho behavior of the flaps may bo gained from figuros 15 and In, which show tho offoct of typical installations on the wing chordwiso prossuro distribution and on tho wing wake at tho horizontal tail piano. Tho drag coefficient duo to all the flaps tosted iß summa- rized in figuro 17 for tho purpose of determining their effect on the velocity of tho airplane It should be noted that tho drag

EA.CA EM Ko. A7F09

coefficient of each installation was diTided by the ratio of projected flap frontal area to wing area in order to account for different flap sizes on tho various models.

Flap offcetivonoss.— Although the purpose of the dive-recovery flaps on a particular airplane is to increase its trim lift coefficient, it is most convenient to consider flap effectiveness as the increase of pitching-momont coefficient at a givon lift coefficient. By so doing, the effect of contor-of gravity position is eliminated. Furthermore, in order to reduce tho data to a form suitable for gonoral application, it is necessary to consider not the total pitching moment increment but tho individual contributing factors. Thoso aro (l) tho offect on the wing pitching-moment characteristics, and (2) tho effect on tho tail lift (or pitching moment due to the tail). Tho socond factor may bo attributed mainly to a change of tail angle of attack and a change of elevator floating angle. (The results indicato that changes of tail efficiency are sma.11 and may be neglected.) Therefore tho total pitching-moment coefficient duo to the flaps may be represented, by the following equation:

In turn, the change of tail angle of attack may be attributed to a change of airplane angle of attack for a givon lift coefficient and to a chango of downwash at the tail due to tho altorod spanwise distribution of lift with tho flaps deflected. Both are dependent upon tho lift developed by the flaps. Hence, the change of tail anglo of attack may be represented by the product of (l) the change of airplane angle of attack for a givon lift coefficient, which is largoly a function of tho size and chordwiso location of tho flaps, and (2) tho ratio of change of tail angle of attack to chango of airplano angle of attack, which is mainly a function of spanwiso location of tho flapa. Thoreforo,

**-«*-* M-SKsS. + *•- d^ (1)

Also, if the elevator characteristics aro linear within the range being considered, and it is assumed that the chango of elevator floating angle ia attributable mainly to the change of tail angle of attack,

NACA EM Ho. A7F09

A&oo =

<&fc

and

ACm = ACn^ + £a (*£ VAa ')G (2)

It should be remombored that the change of tail anglo of attack considered höre ia an average along tho span of tho tail, and equation (2) is applicable only if tho tail characteristics are essentially constant along tho span.

Values of Apm^, £®*x and Axt/Ax were computed for oach flap installation tested and aro givon in figures 18 to 21, since all other factors in th* foregoing equation aro characteristics of tho particular airplane and not of tho flaps. Since most of tho flaps woro tostod at soveral deflections, plots were first made of Ax against the ratio of projected flap frontal area to wing area and of ACm^. against tho ratio of the product of the projected flap frontal aroa and wing chord at tho flap to tho product of wing aroa and moan aerodynamic chord (such as the example in fig. 18). An average linear variation throughout the usable range was assumed as indicated by the dashed linos in the oxample; the slope of these linos was then takon as tho effectiveness of the flaps and plotted against Mach nurabor in figure 19. Values are shown for lift coefficients of 0.00 and 0A0, corresponding to typical conditions for a vortical divo and for moderate recovery from a high-speed dive. In figure 20 tho offoctivonoss is shown as a function of chordwise location of tho flaps for constant values of Mach number relative to the Mach number of lift divergence, However, it should be remembered that tho results are for models having quite varied configurations, and the effects shown may not be due entirely to chordwise location. Comploto data for flaps at various chordwise positions on tho same model are availablo only for the YP—80A model, and the curves shown in figure 20 are faired through the data obtained for this model.

NACA EM Wo. A7F09

Figuro 21 shows a front view of each installation tested and the corresponding values of £a-t/£&,- (Plots wore first made of *at/£a. which rovoaled no approciable variation with flap deflection.) Figuros 22 and 23 show typical offocts of flaps on the elevator characteristics. Figure 2k presents the results of flight teats made with a small airfoil mounted in the high-speed-flow rogion of the airplane wing in order to obtain qualitative information concerning tho effectiveness of dive—rocovory flaps at Mach numbers noar 1.0.

Although it is realized that many secondary factors (such as airfoil section, wing aspect ratio, and support interference), influ- ence the effoctivoncss, it is impossible to complotely determine their magnitude from tho data at hand. For this reason, it is dosirablo that each now installation bo carefully fli^it-tcstod before final approval. Figuro 25 shows the results of such a tost.

DISCUSSION

General Behavior of Dive-Recovery Flaps

Flan deflection.— All the flaps tested were well forward of the wing trailing edge and exhibited tho general characteristics of spoilers. For low Mach numbors the small deflections wore relatively ineffective and in some coses had reversed effectiveness. For high Mach numbers, however, tho reversal tended to disappear, and for practical purposos tho effectiveness may bo assumed to vary linearly with tho proJoeted flap frontal area. (See fig. 18.)

Mach number.— In general, tho effectiveness of the flaps incroased considerably with increasing Mach number to well paBt tho Mach number of lift divergence. (See fig. 19.) As shown in the typical prossuro distribution (fig. 15) the pressure recovery aft of tho flap was less completo at high Mach numbers, causing a considerable incromont of negative pressure ovor a large portion of the upper wing surface. Also, the upper—surface shock moved aft when the flaps wore deflected. Both of theso effects contributed to the flap effectiveness. However, thoro is ovidenco that the effectiveness will docroase sharply abovo somo Mach number between that for lift divergence and a Mich number of 1.0. (See fig. 2k-.) Moroovor, it is bolioved that tho Mach number at which the flap effectiveness decreases will more closely approach the Mach number of lift divorgence as the latter approaches a value of 1.0. It is important, then, that extreme caution bo exorcised in testing dive— rocovory flaps on airplanes having a very high Mach number of lift

8 NACA EM No. A7F09

divergence.

It will bo notod that tho incromont of wing pitching-nomont coofficiont duo to tho flaps became moro negative oe tho Mach numbor increased. (Soo fig. 19.) This is due to .the fact that tho increase of Mach nurabor caused an incroaso of flap effectiveness which was applied largely to tho wing aft of tho position of tho uppor—surfaco shock. In ordor to provont serious changes of tail loads and a docroaso of tho total flap effectiveness, it is dosirablo that tho chango of wing pitching-moment coofficiont ho as small as possiblo.

Location on tho wing.— In presenting the results it was assumed that tho factors Aa. and £Cmw ' are predominantly affected by tho size and chordwiso location of tho flaps on tho wing, while tho factor Ao>t/Ax is largoly a function of tho spanwiso location of tho flaps. This assumption is based on simple thoory and should bo valid except for casos where three-dimensional effects aro largo (such as for flaps noar the wing tip), or for installations which are greatly affected by interference of fuselage, nacollos, etc. Moreover, it should be noted that most of tho wings tosted wero essentially unswept and the values of &D&V given in this report will not necessarily be correct for wings with considerable sweep. A swopt-back wing with flaps inboard of tho moan aerodynamic chord probably would exhibit moro positive values of AQnv Also, tho offectivonoss of tho flapB probably would bo loss on highly swept wings duo to the cross flow.

In general, the flaps locatod well forward on tho chord of tho wing wero moro satisfactory than those noaror the trailing edge. (Soo fig. 20.) Tho forward flaps produced a greater docroaso of anglo of attack for constant lift coefficient, and also caused smaller.negative shifts of tho wing pitching-mamont coofficiont. Largo negative shifts of the wing pitching-moment coefficient should bo avoided since they not only reduce tho total effectiveness of the flaps but may cause serious incroasos of tail loads. However, thoro is a practical limit to tho forward location of the flaps. The range of positive lift coofficionts for which tho forward flaps wore effective was considerably smaller than that for tho roarward flaps. (See fig. 5.) Also, tho flaps in the more rearward positions maintained their offoctivanosa to a slightly highor Mach numbor than did those noaror tho loading odge, especially at tho highor lift, coefficients. Apparently the flow behind tho flaps -which wero woll ahoad of tho lowor-ßurface minimum—pressure point had a strong tend— oncy to return to tho wing surfaco. For this reason, dive—recovery flaps on airplanes requiring a high lift coofficiont for dive rocov— ory should bo locatod farthor aft than those on airplonos requiring

HACA EM Hb. A73T09

a lower lift coefficient. It appears that tho optimum chordwise flap location for airplanes requiring approximately O.lj-O lift coefficient is about one—third of tho wing chord hack of -tee leading edge..

Tho change of tail angle of attack £pt/*a Is shown in f iguro 21 for all the flaps tested. Values of tet^faa, varied from about 0.8 for a model with tho flaps completely outboard of the tell to about 2.0 for a tvin-fusolage nodel with the flaps ontirely in front of The tail. Typical valuoa for flaps partially in front of tho tail woro between 0.9 end 1.2.

Wing soction.— No comprehensive results are available which will show tho offocts of wing soction on the behavior of dive- recovery flaps. It has alroady boen mentioned that the chordwise position of the upper—surface shock might affoct tho wing pitching— moment coefficient duo to the flaps, end that the chordwiso position of the lower—surface minimum-pressure point probably affects tho behavior of flaps located near tho wing leading edge. Within tho range used on airplanes at present., wing—thickness ratio must be accounted for inasmuch as it affects the Mach number of lift divergence.

Elevator characteristics.— In developing equation (2) of thö section ontitled "Sesults," It was assumed that the change of elevator floating angle due to the flaps arose from the change of tail angle of attack. This assumption is substantiated by the rosults shown in figures 22 ejai. 23. Tho dive—recovery flaps did not greatly chango tho elevator floating angle of tho model whose elevator hinge moments wore essontlally unaffected by changes of tail angle of attack (or airplane lift coefficient). However, a large chango of elevator floating anglo was noted for the model the elevator hingo moments of which varied considerably with tail angle of attack. At a Mach numbor of 0.70, tho elevator hinge moments with tho flaps deflected appoared to be about the same for a lift coofficiont of Q.hO as they wore with the flaps retracted for a lift coefficiont of about —0.1. This indicates that tho flaps decreased tho tail anglo of attack about 3«5° (using a measured value of 7.0° for bat/bCjJ. 5310 lift and pitchIng-noment results (figs. 11, 19, and 21) indicate a decroaso of tail angle of attack of 3.1° and 3.6° at lift coefficients of 0.00 and 0.1{-0, respectively. In view of tho absence of inforaation for a wider range of models, the results are conclusive enough to .Justify the assumption.

It is significant that tho flaps did not change tho elevator effectiveness of the two models for which data are available. As a result, it can be assumed that the horizontal—tail effectiveness

10 HACA EM No. A7F09

would not bo seriously affected, end flap—retracted values of (dCn/döoJat aaö- (dCm/dat)a ^^ be usod in the oquations for total p itching-moment coefficient due -to the flaps.

Structural and Mechanical Considerations

Flap loads.— Limited prossuro-niistribution noasurononts have boon made on typical divo—rocovery flaps which indicate a trapezoidal chordwise loading with a maximum loading at tho hinge lino. For practical design purposes a uniform chordwiso loading may bo usod. Except for snail dofloctions where the pressures night be reversed, tho normal—force coefficient may bo assumed to- vary linearly with flap defloction to a value of 1.1 for a deflection of h^°.

Kate of flop deflection.— The flap actuating mechanism should be supplied with sufficient power to deflect the flaps rapidly. Tho rates of flap defloction for successful installations tested in flight have boon such that full flap deflection was readied in 1 to lias- seconds. Slower rates of deflection result in dangerous response lag and a consequent greater loss of altitudo during recovery from a divo. In .addition, tho adverse offoots of the flaps for small deflections are accentuated.

Buffeting.— It has boon shown that the most effoctivo spanwiso location for dive—recovery flaps is directly in front of tho horizontal tail. Although no conclusive data are available on tho matter, there may "bo some danger of tail buffeting with flaps in this location. Therefore, a compromise spanwiso location is roconraondod that places the flaps in front of only tho outboard portions of the tail, especially on airplanes which do not have a particularly high tail position with rospect to tho wing. Also, isolated cases of wing buffeting have been roportod with divo— rocovery flaps deflectod, wherein the landing flaps vibrated through a comparatively small amplitude. It is bolievod that this was duo to a small amount of play in tho landing—flap rostraint mechanism.

Recommended Design ProcoduTo

It has been shown that the effectiveness of divo—rocovery flaps in increasing the pitching-momont coofficiont of an. airplano is maintained past tho Mach number of lift divergence for the airplano. However, the trim lift coefficient and resultant acceleration ero loss at Mach numbors above that for lift divergence bocauso of tho increasod static longitudinal stability. (Seo fig. 25.) -Hence, a

NACA EM No. A7F09 11

successful dive—recovery flap Installation is one •which, will effect sufficient (hut not excessive) recovery of the airplane from a dive to any Mach number and altitude, which the airplane is capable of attaining, without causing undue stress on any part of the aircraft structure. The following procedure for the preliminary design of dive—recovery flaps for a particular airplane is offered:

1. Establish the Mach number, altitude, and lift coefficient at which the recovery is desired. (3he drag due to the flaps may be neglected since appreciable changes of drag affect the velocity only slightly above the Mach number of lift divergence.)

2. Compute or estimate the elevator-free pitching-moment characteristics of the airplane with the normal center—of—gravity position for the above Mach number." From those characteristics determine the pitching-moment—coefficient, Increment necessary to trim the airplane at the lift coefficient of step 1.

3. Select a flap location on the wing which is structurally suitable for the airplane (around 30 or kC percent of the chord and partially In front of the horizontal tail if possible).

4. Assume the flap size for a comparable installation from figures 3 to lk.

5. Compute valuos for t£L, £Cmw* 8^- ^at/Ax from figures 20 and 21 "for the flap size assumed and a deflection of 30°. (ibis will allow some adjustment of the effectiveness during flight tests if necessary.)

6. Compute £Cm using the equations developed in the section entitled "Besults." If the elevator characteristics are linear within the range being considered, equation (2) may be used; other- wise, the change of elevator floating angle should be determined directly from the elevator characteristics and equation (l) should be used.

7. If the £Cm computed In step 6 does not agree with that of step 2, repeat the procedure with a different flap size.

Once the flap Bize is established, It would be wise to compute the acceleration available from the flap3 for all Mach numbers and altitudes which the airplane is capable of attaining, for both the most forward and the most rearward center—of—gravity position. Also, tho tail loads with the flaps deflected should be checked if it

12 HACA EM Wo. A7F09

appears that the flap location chosen will result in an appreciable value of ÜCvfy.

Recommondod Flight-Tost Procedure

In viov of previously mentioned deficiencies of the available data, it is desirable that each new dive—recovery—flap installation bo carefully flight-tested before final approval. A stop-by—stop procedure is recommended which consists of trimming tho airplano for a given spood and deflecting the flaps. Tho stick—froo condi- tion should bo simulated and tho olovator control used only to prevent excessive accelerations. Measurements should bo made of tho Mach numbor and altitudo at which idle flaps are deflected and of tho maximum accoloration obtained during tho rosulting maneuver.

Runs should be made throughout tho speed range, but it is desirable that tho first runs be made at comparatively low Mach numbers and high altitude in order to avoid excessive accelerations. A running record should bo kept so that the flap effectiveness at the higher Mach numbers may bo anticipated.

Particular care should bo exercised in obtaining data above tho Mach number of lift divergence for tho airplane. The desirod Mach number should bo approachod gradually at the minimum diving angle necessary to attain that Mach number at the altitude being used. As a result, the maximum Mach numbor of any particular run will be only slightly greater than tho Mach number for which data have alroady boon obtained. Following tho ahove precautions will rosult in obtaining the necessary information with a maximum degree of safety.

CONCLUSIONS

An analysis of the data colloctod in this report indicates the following:

1. In order to obtain maximum effectiveness, as well as to avoid tail buffeting and largo increasos of tail loads, divo— recovery flaps should be located about one—third of -the wing chord from tho loading edge with part of the flap in front of the outboard portion of the horizontal tail.

2. All tho flaps testod were effective above the Mach numbor of lift divergonco for tho airplano.

NACA EM No. A7F09 13

3. Tho flap offectivenose probably will become negligible at some Mach number between that of lift divergence for the airplane and a Mach number of 1.0.

h. All new dive—recovery-flap installations should be carefully flight—tested, especially "chose intended for use on airplanes with a. very high Mach number of lift divergence.

Ames Aeronautical Laboratory, National Advisory Committee for Aeronautics,

Moffott Field, Calif.

o >

>

8

Figure 1.- Typical dive-recovery-flap installation on a wind-tunnel model.

£

NACA RM No. A7F09 15

(a) Flap deflected 30°.

(b) Flap fully retracted.

Figure 2.- Dive-recovery flap located on left wing of the XP-51 airplane.


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r TITLE: A Summary and Analysis of Data on Dive-Recovery Flaps

AUTHOR|S)t Boddy, Lee E.; Williams, W. G. ORIGINATING AGENCY: National Advisory Committee for Aeronautics, Washington, D. C PUBLISHED BY; (Same)

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ATI- 12043

(None)

HltUlHING AGENCY NO

"

OA11

Sept'47 ABSTRACT:

U.S. gne. KUHTXATIONI

phnt >K, graphs, drwys

Results of numerous unrelated tests of dive-recovery flaps are collected and presented In a form suitable for use in preliminary design of dive-recovery flap installations. Since data were obtained for airplane models of widely varying configurations and are limited largely to Mach 0.80, it is recommended that each new installation be carefully flight-tested before final approval. A flight-test procedure is outlined which will insure a maximum degree of safety.

DISTRIBUTION: Request copies of this report only from Originating Agency DIVISION: Aerodynamics (2) SECTION: control Surfaces

ATI SHEET NO.: R-2-3-&4 TWJ SUBJECT HEADINGS: Flaps, Dive (37455); Flight testing - Airplanes (38425)

Air Documents Division, Intolligcftc* Department Air Material Command

AIR TECHNICAL INDEX Wrlght-Pattvrso« Air Fore* I Dayton, Ohto

i;t •

UNCLASSIFIED per authority NACA Notice of De clarifi- cation of Publications No. k, <itd April-Septetnhar 195c.

'//--da/* 7