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NATIONAL ADVISORY ,COMMJITEE ~ :. ~ ... -:;.-

FOR AERONAUTICS

TECHNICAL NOTE

No. 1668

INVESTIGATION OF EFFECTS OF GEOMETRlC DIHEDRAL ON LOW-8PEED

STATIC STABILITY AND YAWING CHARACTERlSTICS OF AN UNTAPERED

45° SWEPT BACK-WI~"G MODEL OF ASPECT RA.TIO 2.61 .• -:- .. •"'··'- .::..-- -. :,,t-

By M: J. Queijo and Byron M. J~quet

.I:.angley Aeronautical Laboratory Langley Field, Va.

~nashington

September 1948

NATIONAL ADVISORY COMtviTTrEE FOR AERONAUTICS

TECHNICAL NOTE NO. 1668

IN'VE3'l'IGATION OF EFFECTS OF GEOMETRIC DIHEDRAL ON WIT -BPEI!:D

STATIC STABILITY AND YAWING CHABACTERISTICS OF AN UNTAPERED

45° SWEPI'BACK-WING MODEL OF ASPECT RATIO 2.61

By M. J. Queijo and Byron M. Jaquet

SUMMARY

An investigation was conducted to deter.mine the effects of geometric dihedral on the·low-speed static stability and yawing characteristics of an untapered 45° sweptback-wing model of aspect ratio 2 .61. The results of the tests indicated that an increase in positive dihedral resulted in an increase in the rolling mo.:nent due to sideslip and also caused the maximum value of rolling moment due to sideslip to occur at increas­ingly higher lift coefficients. _Increasing positive or negative dihedral caused a decrease in the lift-curve slope and an increase in the variation of lateral force with sideslip. DLhedral had no appreciable effect an the yawing moment due to sideslip • ·

The rolling moment due to yawing became more positive with increas­ingly positive dLhedral and became less ~ositive with increas~ngly negative dihedral. The rate of change of rolling moment due to yawing with dihedral angle was nearly independent of' lift coefficient. The yawing moment due to yawing was nearly independent of lif't coefficient for low and moderate lift coefficients and showed no definite trends at higher lift coefficients. The lateral force due to yawing became more positive with an increase in positive or negative dihedral and sho¥ed little variation With lift coefficient through the low and m.ode~te range of lift coefficients. At higher lift coefficients, the lateral force due to yawing became more positive.

INTRODUCTION

Estimation of the dynamic flight ·characteristics or airplanes requires ~knowledge of the component forces and moments resulting from the orientation of the airplane with respect to the air stream ~d from the angular velocity of the airplane about each of its three axes. The forces and moments resulting from _the orientation of the airplane usually are expressed as the static stability derivatives, which are readily determined in conventional wind-turmel tests. The forces and m:.>mento related to the angular motions (rotary derivatives) generally have been estimated. from theory because of the lack of a convenient experimental technique.

2 MCA TN No. 1668

The recent application of the rolling-flow and curvea-flow princi­ple of the Langley stability tunnel has made possible the deter.mination of both the rotary and stat~c-stability derivatives with about the same ease. Unpublished data have indicated that althougn the rotary stability deri va ti vee of unswept wings of --m.od.e:ra te or high aspect . ratio can be predicted quite accurately from the available theory, the use of sweep -and,perhaps, low aspect ratio - introduces effects which are not readily amenable to theoretical treatment-. For this reason, a systematic research program has been established for the purpose of determining the effects of various geometric variables on both rotary and static stability characteristics.

The present investigation is concerned with the deter.m.ination of the effects of geometric dihedral an the static stabi~ty and yawing characteristics of an untapered 45° swept wing of aspect ratio 2.61.

SYMBOLS

All forces and moments are given with respect to . the stability axes with the origl.n at the quarter-chord point- of the mean aerodynamic cho:rd of the wing. The poai tive direction of-the forces, momenta, angular displacements, and velocities are shown in figure 1. The symbols and coefficients used herein are defined as follows:

CL lift coefficient (L/qS)

CL I lift coefficient based on lift of ons panel o.f wing with dihedral and on area of entire wing

Cy lateral-force coefficient (Y/qS)

Cx longi tudinal-i'-Orce coefficient (X/ qS)

C1 rolling-moment coefficient (L/qSb)

Cn yawing-moment coefficient (N/qSb)

Cm pitching-moment coefficient (M/qSc)

L lift, pounds

Y lateral force, pounds

X longitudinal force, pounds

L rolling moment about x-axis, foot-pounds

N yawing moment about Z-axis, foot "":Pounds

M pitching moment about Y-axis, foot~ounds

NACA TN No • l668

q

p

v

s

dynamic pressure, pounds per square f'oot (~v 0 2)

mass density o,t' air, slugs per cubic f'oot

f'ree-streamvelocity, f'eet per second

local velocity, f'eet per second

wing area (zero dihed.raJ. wing) , square f'eet

3

b span of' wing, measured per:pendicular to plane of' symmetry (zero dihedral wing), f'eet ·

c

y

X

chord of' wing, measured parallel to plane of' symmetry, f'eet

absolute value of' spanwise distance f'rom plane of' Syill!D.etry to e:n:y station on Wing quarter-chord line

mean ae~c chord, feet (~ [12

c2 d0 rearward distance f'ram. coordinate origin (airplane center of'

gravity) to aerodynamic center

ef'f'ective lateral center-of-pressure location of' the resultant load caused by rolling

A aspect ratio (b2js)

a. angle of' attack measured in a vertical plane parallel to the plane of' symmetry, degrees

t angle of' yaw (equal to -~), degrees

13 angle of' sideslip, degrees, unless otherwise indicated

A angle of' sweep, positive f'or sweep back, degrees

r dihedral angle' degrees (unless otherwise specified)

rb/2V 0

yaWing-velocity parameter

r angular velocity in yaw, radians per second

R radius of' curvature of' f'llgb.t :path

a 0 section llf't -cUTV"e slope, per radian

4

deL Crc. =­Cia.

den

Cn"' = d'if

dey

Cy V =. Clv

dey Cyr =

21rb

2Vo

NACA TN No. 1668

dczv/Cir dihedral-effectiveness parameter; rate of ch~ of Czv with dihedral angle

dezr/Cir rate of change of Czr with dihedral angle

Subscripts :

i induced

L left "'Wing pa.riel

R right-wing panel

APPARATUS AND TESTS

The teats of the present investigation were made in the 6- by 6-foot teat section of the Langley stability tunnel, in which curved f'low can be ~

simulated by curving the air stream about-a stationar;y_model.

NACA TN No . 1668 5

The model tested was an untapered 45° sweptback wing (see fig. 2) with a 10-inch chord and NACA 0012 airfoil sections in planes normal to the leading edge. The model was constructed of laminated mahogany and consisted of two panels which were joined together by metal brackets. The brackets were made to give the model dihedral angles of 10°, 0°, -10°, and -200. The model was rigidly attached to a single strut into which was built a strain-gage balance system by which all the forces and moments on the model were measured. A photograph of the model mounted on the support strut in the curved-flow test section. is shown as figure 3 · Some clearance was provided between the· strut and the model. No attem;pt was made to seal the clearance gap because previous tests of a similar model showed that sealing the gap had negligible effect an the characteristiss of the model. ·

Two series of tests were made. The first series consisted of straight-flow tests for each model configuration. The tests were made through a yaw-angle range fran -30° to 30° for several angles of attack. The second series of tests were made in simulated yawing flight for each model configuration (r= 10°, oo, -10°, and -20°) . The yawing-flow tests were made at zero yaw angle and at stream curvatures corresponding to values of rb/2V 0 of 0, -o .031, -o .067, and -o .o88, based on the span of the zero-dihedral model. Each model configuration was tested fran about z~ro lift up to the stall.

All teats ware made at a dynamic pressure of 24 ·9 pounds per square foot, which corresponds to a Mach number of 0 .13 and a

6Reynolds number,

based on the model mean aerod.ynamic chord., of l.l x 10 .

CORRECTIONS

Approximate· correcti.ons for jet -boundary effect were applied to the angle of attack and to the longitu~l-force coefficient. A correction was also applied to the lateral force to account for the error caused by the static -pressure gradient across the curved -flow test section.· The corrections used are:

.6Cy = 4 .o..:L rb bS 2V

6 NACA TN No . 1668

where

6w boundary correction factor from reference l

S' tunnel cross-section area at test section, feet

v volume of model, cubic feet

No corrections were made for tunnel blocking or eupporl strut tares except for the derivatives C~ • In this case, the tare at-zero lift coefficient r was applied throughout the lift-coefficient range.

RESULTS AND DISCUSSION

Presentation of Data

All. the test data are based on the area, span, and mean aerod.ynamic chord of the zero-dihedral model configuration. The data obtained from the straight -flow tests are presented in figures 4 and 5. Curves Qf lift coefficient Or,, longitudinal-force coefficient Ox, and pitching­moment coefficient Cm plotted against angle of attack for each mod.el . configuration (t = 0°) are presented in figure 4. Curves of rolling­moment coefficient C~, yawing-moment coefficient Cn, and lateral­force coefficient Cy plotted against angle of yaw for several angles of attack and for each model configuration are shown in figure 5. The data obtained from the yawing-flow teats are presented in figure 6 as plots of C 1. , Cn, and Cy against rb /'ZV 0 for several angles of attack for each model configuration.

The variations of CZ-f, ~, and Cy'lt with lift coefficient are

shown in figure 1 for each model configuration in straight flow. The variations of Clr' Cnr' and CYr with lift coefficient are shown in figure 8 for each model configuration.

The effects of dihedral on the static longi tudina.l stability character­istics Cr.a, and d.Cm/d.CL are shown in figure 9. The effect of dihedral on the rolling-moment derivatives C7,t and C7.r for several lift coeffi­

cients is shown in figure 10. The variations of the parameters ?IJ7,v/Or n.nd ?i:Jlr/'dr with lift coefficient are shown in figure ll·

Straight-Flow Results

Lift characteristics.- The slopes of the lift curves at ·zero lift of figure 4 are presented. as a function of dih~dral angle in figure 9 • The curve of figure 9 indicates that the lift-curve slope decreases with increasing positive or negative dihedral. The analysis of reference 2

NACA TN No • 1668

shows that the variation of ~ift -curve slope with dihedra~ can be expressed as

7

(~)

where ( CLcx. )r=oo is the lift -curv-e s~ope o:f the zero-dihedra~ mode~ and

(era. )r is the lift-curve. s~ope of the same mode~ With dihedra~. The

curve of (era. )r obtained by means of' equation (~) and the measured

va~ue of (Cia, )r=Oo are presented in figure 9 for comparison With the

experimentally obtained variatio.::~.. The two curves genera~ are in good agreement and are consistent with simi~ar re~ts in reference 3·

Pitching-moment characteristics.- The pi tching-m.oment data of figure 4 indicate that a& the dihedra~ angle is mad.e positive, the pitching moment geneZ'8J.l.y becomes ~ess positive • The s~opes of the pitching-moment curves of figu.re 4 were measured at zero ~1ft and are plotted in figure 9 as a curve of OCm/dcr. against dihedral angle • The slope dCm/~L generally becomes sligb:tly ~ess positive as the dihedral angle is made more positive .

Longitudinal-force characteristics.- The longitudinal force is nearly independent of dihedra~ angle for angles of attack up to about 12°. (See fig. 4.) At higher angles of attack the ~ongitudina~-force coeffi­cient generally decreases with either positive or negative dihedral.

Rollip.g-mament characteristics.- The rolling-moment data of figure 7 indicate that, for the wing tested, positive dihedral resulted in a positive displacement of the curve of C~t and negative dihedral resulted in a negative displacement of the curve of C~t at all ~ift

coefficients • As the dihedral angle becomes more positive, the :ma.ximum. value of 0 z t occurs at increasingly higher lift coefficients; whereas, increasing the dihedra~ negatively causes the maximum value of Czv to

occur at increasingly lower lift coefficients. This trend is exacUy the opposite to that reported in reference 3. The disagreement is believed to be caused by the differences in taper ratio and in camber of the two mode~. The model of reference 3 had a taper ratio 0 ·5 and a Bhode. St. Genese 33 airfoil section.

The rolling-moment data of figure 7 were cross-plotted in figure ~0 to give curves of C~'t as a function of dihedral angle for several lift

coefficients. The curves of figure 10 indicate that the slope of the curve of' 0~ 'ljr against dihedral angle generally is constant in the -10° to 10° dihedral-angle range and decreases slightly in the -10° to -20° range • The same trend was indicated in reference 3. The slopes of ~e

8 _ NACA TN No . 1668

curves of Cz t of figure 10 were measured in the -100 to 100 dihedral­

angle range and were plotted against lift coefficient in figure ll· The plot indicates that dc7.~dr is nearly independent of lift coefficient

up to a lift coefficient of about 0 ·5 and has a value--of about 0 .ooou. At higher lift coefficients dczw/dr increases to 0.00017·

The curve of dc7.t/dr of reference 3 is included in figure ll for

comparison with the results of the present inv-estigation. It is seen that the curves of the two investigations are quite different both in magnitude and in mode of variation with lift coefficient. In an attem;pt to explain these differences, an equation based on the methods of refer­ence 4 and e:x:tended to include dihedral angle effects was derived. (See appendix.} The equation is

dczw --= dr (A + 4) cos A (dczw \ A + 4 cos A -g:r-k.=0 (2)

( doz-.\ where ~ )ll;:::.()o is the dihedral effectiveness parameter o:r an unewept

Wing with the same aspect ratio as the Wing_'llD.der cone:J,.derat"ion and is obtained directly from figure 12 for aspect' ratios fro:n l to 16 and taper

· - - d0z ratios of 1.0 and 0.5. Equation (2) gives a result of ~ = 0.00013

dr for the wing of the present inv-estigation and a result-_ of 0.000143 for the wing of reference 3. Although these calculated results are not very good checks on the experimental data, they do show that equation (2) gives the proper trends as far as magnitude is concerned. Little more may be said for the validity of equation ( 2) until more data are available for comparison with calculat~d results.

Yawing-moment characteristics.- The yawing-moment data ofu figure 7 indicate that C~ is nearly independent of dihedral for lift- coefficients

up to approximately CL = 0 ·5. At higher ·lift coefficients the curves of Cnt are irregular, but in general, Cn;r becomes more positive With

increasingly negative dihedral and more negative with increasingly positive dihedral.

Lateral-force charact-eristics • - The lateral force due to yaw Cyt

becomes more positive as the dihedral is made more positive or more negative. Cyt is nearly independent- of lift coefficient up to about

CL = 0 ·5 and becomes irregular at higher lift coeffic_ients. _ (See fig. 7 .) _

NACA TN No o 1668 9

Yawing Flow

Rollinp:-mament characteristics o- The rolling-moment data of :figure 8 indicate that the rolling moment due to yawing becomes more positive with increase in positive dihedral. Reference 5 indicates that the dynamic stability of an airplane decreases ~th a decrease in positive C?.r• This result indicates that the use of negative dihedral might decrease the dynamic stability; however, whether negative dihedral is detrimental or beneficial to dynamic stability depends an the effect of dihedral angle on all the derivatives which affect dyna.mic stability. Figure 8 also indicates that Clr generally increases with lift coefficient over the low-li:ft-coef':ficient range.

The rolling-moment data of' :figure 8 were plotted in :figure 10 as curves of C ?.r against dihedral :for several lift coefficients o The

derivative C7.r varies approximately linearly with dihedral :for a given lift coefficient • The slopes of the curves of :figure lO were measured and plotted in :figure ll as a curve of' Oc?.r/or against lift coefficient. The parameter OC?.r/or is practically independent of lift coefficient and has an average value of about 0 .oo4o. The methods of' reference 4 were extended to include the effects of small amounts of dihedral, and the :follO"tTing equation was deriv~ (see appendix) for a sweptback wing with dihedral :

Oczr 1 1CA sin A ar = .i2 A + 4 cos A

where r is in radians • For the wing used in this investigation, Ocz

equation ( 3) gives a value ---=:.. = 0 .o~ and converting r to degrees or OC?.r

gi vee a value ~ = 0 .0016. This value is less than half' of the

(3)

average value ( 0 o0040) obtained in the tests reported herein. Although equation (3) indicates the proper trends of the effect of' sweep on the para.me:ter OC?.r/Or, the magnitude of the effect is much too low. It is possible that the value Oczr/?JI' giv~n by equation (3) should be consid­ered sim!Jly as an increment due to sweep and that it should be added to the value of Oczr/Or' of' unswept wings in order to get the total value :for a swept Wing. (Such was :found to be the case in reference 4 for the derivative Cz~ of' ~ept Wings.) Whether this hypothesis is correct cannot be determXned at this time because of the lack of data on the d.erivative dcLr/OI' of unswept and swept Wings.

Yawing-moment characteristics.- The yawing-moment data of figure 8 indicate that Cnr is nearly independent of dihedral and also of'

10 NACA TN No. 1668

lift coefficient up to a lift coefficient of ab~t CL = 0·5·, The curves are irregular at higher lift co3ff1cients.

Lateral-force characteristics.- The lateral-force data of figure 8 indicate that increasing the positive or negative dihedral generally :makes Cyr leas negative and that Cyr varies only slightly with lift coefficient.

CONCLUSIONS

An investigation was conducted in the 6- by 6-foot teat section of the Langley stability tunnel to determine the Bffects of dihedral on the aerodyna.mic characteristics of an untapered 45° sweptback""Wing model of aspect ratio 2.61 in atraight-f'low and in yawing flow. The results of the investigation have led to the following conclusions :

l • The results obtained for the low -speed eta tic stabill ty charac­teristica were generally consistent With those o£ previous investigations.

An increase in positive dihedral resulted. in an increase in the _ :~oiling moment due to sideslip for all lift coefficients and also caused ·i ... .!le maximum value of rolling moment due to sideslip to occur at increas­ingly higher lift- coefficients •

Increasing positive_ or negative dihedral caused a decrease in the lift -curve slope and an increase in -the variation of lateral force with sideslip.

Dihedral had no appreciable effect on the yawing moment due to sideslip. --

2 • The rolling moment due _to yawing became more positive w1 th increasingly positive dihedral and became less positive with increasingly negative dihedral.

3. The rate of change of rolling moment due to yawing with dihedral angle.was practically independent of lift coefficient and had a value of about o.oo4o per degree of dihedral.

4 • The yawing moment due to yawing was nearly independent of dihedral angle and lift coefficient for lift coefficients up to-_about~O .5 but showed no definite trend at higber.lift coe~cients.

NA.CA TN No • l668 ll

5. The lateral force· due to yawing became more positive with increase in positive or negative dihedral and showed little variation with lift coefficient.

Langley Aeronautical Laboratory National Advisory Committee for Aeronautics

Langley Field, Va., April l4, l948

12 NACA TN No. l668

APPENDIX

Approximate equations were derived in re~erence 4 for the stability derivatives o~ swept wings without dihedral. The methods of re~erence 4 are extended herein to evaluate the parameters dcz~/or and dczr/or.

Dihedral Effectiveness Parameter

For wings with dihedral tlie change ilL angle of attack-~ resulting from aide slip, can be shown to be L:n = ~ sin r, where lY:J. is measured in planes pe:r:pendicular to each wing panel and parallel to the relative wind. For an antis;ymm.etrical load distribution the induced angle in the

14CL' same planes is appro.ximately b:r-i = ~· The li~t--curve slope of'--an

infinite skewed wing is ao cos A; therefore, f'rom. lifting-line theory

(Al)

Substitution of' the values of' lY:J. and ~ into equation (Al) gives

or

1 Aa0 cos A 0L I = 2 2ao cos A ~ sin r

A+ '1t

If 2'1t is substituted in the denominator f'or a 0 , then

-1 __ Aa0 cos A a

CL' ~ ~ sin r - A + 4 cos A (A2)

NACA TN No . 1668 13

The coefficient CL' is of the same magpitude but of opposite sign for the two wing panels. The rolling-moment coefficient for the entire wing therefore is given by

Substitution of the value of CL' given by equation (A2) into equation (A3) results in

Aaa cos A v c1 = - ~ sin r L A+ 4 cos A b

If r is small, then sin r = r, and

Cz j3 l Aa0 COB A _,i_

r = -2 A + 4 cos A b/2

where r is in radians. For unawept wings, equation (A4) gives

(A3)

(A4)

If the approximation is made that sweep has no effect on y, then equations (A4) and (A5) may be combined to give the following equation:

C'Z13 = (A + 4) cos A (Cz~ \ (A6) r A + 4 cos A r ~=00

I'b.e values of (C'Zj3 \ given in figure 16 of reference 6 for r )h=Do

aspect ratios 6 to 16 and taper ratios 1.0 and 0.5 have been extrapolated to low aspect ratios by the procedure used in reference 4. The extra­polated curves are presented in figure 12.

14- NACA TN No. 1668

Rate of Change of Rolling Moment Due to

Yawing with Dihedral

A swept wing wi~h dihedral undergoes a change in angle of attack in flight in a curved 'Path. The change in angl-e. of att-ac_J,c can be shown to be given approximately by

b\ -- ij:')tan A+ X - R sin r (A7)

The velocity over any section of the left wing panel is

VL = r(R + y) (A8)

and, for the right wing panel .

VR = r(R • y) (A9)

The part of the wing loa.ding_ caused by the flight -path curvature- is uns;ymnetrical because of the. velocity gradient. across the wing span; therefore, the rate of change of lift on any section with angle .of attack is given approximately by

Aa0 COB A (V _'f A+ 4 COB A ;;-;;

The rolling moment for rectangular wings is

•

NACA TN No. l668 l5

or

(AlO)

By means of equations (A8), (A9), and (AlO) and if r is assumed to be small, it can be shown that

---- -tanA+-C l.r l Aa0 cos A l x ) r - 4 A + 4 cos A 6 b/'2

If x is zero (as in the tests reported herein) and '2rr is substituted for a 0 , then

C Z.r .1... 1!A sin A r = 1.2 A + 4 cos A

(All)

16 NACA TN No . 1668

REFERENCES

l. Silverstein, Abe, and White, James A.: Wind-Tunnel Interf'erence with Particular Reference to Of'f'-Center Positions of' the Wing and to the Down-wash at the Tail. NACA Rep. No. 547, 1935.

2. Purser, Paul E., and Campbell, John p,: Ex.perimental Verif'ication of' a Simpllf'ied Vee-Tail Theory and Analysis of' Available Data on Complete Models with Vee .. Tails. _ NAC~ R~p. No . 82~, 1945.

3. Ma.gg:l..n, Bernard, and Shanks, Robert E.: The Ef'f'eciJ-o:E Geometric Dihedral on the Aerodynamic Characteristics of' a 400 Swept-Back Wing of' Aspect Ratio 3. NACA TN No. 1169 .r 1946.

4. Toll, Thomas A., and Queijo, M. J.: Approximate Relations and Charts f'or Low-speed Stability Deri va ti ves of' Sw~pt Wings. NACA TN No. 1581, 1948.

5· Sternf'ield, Leonard: Same Considerations of' the Lateral Stability of' High-Speed Aircr.af't. NACA TN No. 1282, 1947.

6. Pearson, Henry A., and Jones, Robert--T.: Theoretical Stability and Control Characteristics of' Winss with Various Amounts of Taper and Twist-. NA.CARep. No. 635,1938.

NACA TN No. l668

X---,._j_A

1f -

x--

Wmd d!re-cfton

;z---_ I

~

Wmd dtrect;on

-~

r lz A-A ~

F~gur(; 1.- The .stability system of axes. Arrows ;ndtcate po.stttv~ dtrect;on.s of forces~ moments.) angular veloc!f!e.s, and angular d;spkTcemenfo.

l7

1.8 NACA TN No. 1.668

T1p of revolutton

Agure 2..- D10grom of model te~tec/.

t

·• I·

r . i jf·' i I I r

' l rr ..

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r I

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' .. ' .

'.

..,,_,.., .. ...___~ ---. - NACA -· ~ ... h"

... :..: ... ~ .• L-52405

0 . Figure 3.- Rear view of untapered 45 sweptback-wing model in the 6- by 6-foot curved-flow

test section of the Langley stability tunnel. a. - 10°; r- 10°.

~ ~ §1 . f-'

~ OJ

NACA TN No • 1668

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~ ~ ~ Ill: i».

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TITLE: Investigation of Effects of Geometric Dihedral on Low-Speed Static Stability and Yawing Characteristics of an Untapered 45° Sweptback-Wlng Model of Aspect *

AUTHORS): Gueljo, M. J.; laquet, Byron, M. ORIGINATING AGENCY: Langley Aeronautical Lab., Langley Field, Va. PUBLISHED BY: National Advisory Committee for Aeronautics, Washington, D. C.

ATO- 34510

(None)

rvBUSKtNO AOSMCT KO.

(Same)

Sept' 48 Unclass. U.S. Eng. 29 ILUKT0AT1OM

photo, diagrs. graphs ABSTRACT:

The results of tests Indicated that an increase In positive dihedral resulted in an increase in the rolling moment due to sideslip and also caused the maximum value of rolling moment due to sideslip to occur at increasingly higher lift coefficients. Increasing positive or negative dihedral caused a decrease in the lift-curve slope and an increase in the variation of lateral force with sideslip. Dihedral had no appreciable effect on the yawing moment due to sideslip. The rate of change of rolling moment due to yawing with dihedral angle was nearly independent of the lift coefficient. The lateral force due to yaw- ing became more positive with an increase in positive or negative dihedral. At higher lift coefficients, the lateral force due to yawing became more positive.

* Ratio 2.61

DISTRIBUTION: Request copies of this report only from Publishing Agency DIVISION: Aerodynamics (2) SECTION: stability and Control (1)

ATI SHEET NO.: R-2-1-54

SUBJECT HEADINGS: Wings, Swept - Stability and control (99304.4); Dihedral (30215); Yaw (99800); Airplanes - Stability (08487)

Air Documents Division, rntolligci AIR TECHNICAL INDEX Wrlght-Patlorton Air Porco BOM

Dayton, Ohio

I TITLE: Investigation of Effect of Geometric Dihedral on Low Speed Static Stability and Yaw

ing Characteristics of an Untapered 45° Swept-back- tfing Model of Aspect Ratio AUTHOR(S): Queijo, M. J.; Jaquet, Byron M. (2.61 ORIGINATING AGENCY: Langley Memorial Aeronautical Lab., Langley Field, Va. PUBLISHED BY: National Advisory Committee for Aeronautics, //ashington, D. C.

ATT0- 35059

(None) OMO. AOENCT NO.

TN-1668 PUOU1HINO AGENCY HO.

(Same)

Sent ' 48 Unclass. U.S. Eng. 28 IUUSTIATIONS

photo, graphs, drwg ABSTRACT:

Tests made to determine the effect of dihedral on the static-stability and yawing derivatives of 45° swept - back-wing model of aspect ratio 2.61 were made in a curved-flow section at a Reynolds number of approxi- mately 1.1 x 10° and consisted of force measurements throughout the lift-coefficient range. The results of the tests indicated that an increase in positive dihedral resulted in an increase in the rolling moment due to sideslip and also caused the maximum value of rolling moment due to sideslip to occur at increasing- ly higher lift coefficients. Increasing positive or negative dihedral caused a decrease in the lift-curve slope and an increase in the variation of lateral force with sideslip. Dihedral had no appreciable effect on the yawing moment due to sideslip.

DISTRIBUTION: Request copies of this report only from Publishing Agency DIVISION: Aerodynamics (2) SECTION: vVings and Airfoils (6)

ATI SHEET NO.: R-2-6-79

SUBJECT HEADINGS: vVings, Swept-back - Aerodynamics (99305.2); .v"ings, Swept - Stability and control (99304.4); Dihedral (30215)

Air Document* Divfilon, Intel lig onto Do part mo nt Air Maloriol Command

Ain TECHNICAL INDEX Wright-Pottorson Air Forto Baso Dayton, Ohio