naca tn-1722 power effects

8/13/2019 NACA TN-1722 Power Effects

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.-1 ?,(_//-NATIONAL ADVISORY COMMITTEE

FOR AERONAUTICSTEC.~CAi NOTE

No. 1722

PREDICTION OF THE EFFECTS OF PROPELLER OPERATION ONSTATIC LONGITUDINAL STABILITY OF SINGLE-ENGINETRACTOR MONOPUNES WITH FILMS RETRACTED

~ J oseph Weil and William C. Sleemm, J r.Langley Aeronautical LatmalmyLangley Field, Va.

WashingtonOctober 1948

. II ,3.

THE

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TECHLIBRARYKAFB,NMInlnullllululllllnNATIONAL ADVISORY COMMU5CEE FOR AERONAUHCS 0144904

TECHNICALNOTENOO lq22

PREDICTION OF THE EET ECISF PROPELLER OPERATION ON THESTATIC LONGIT6= STABILITY OF SINGLKENGINE

TRACTOR MONOPLANES WIT EFLAl% RETRACTEDBy Joeeph Weil and William C. Sleeman, Jr.

ISlmmm

The effects of propeller operation on the static longitudinalstabili@ of single-engine tractor monoplanes ae analyzed, and a simplemethml is presented for computing power-on pitching+uoment curves forfla~retracted flight conditions. The methods evolved me based on theresults of powered+nodel winiL-tunnelinvestigations of 28 model configurations. Correlation curves we presented from which the effects ofpower on the downwash over the tail and the stabilizer effectiveness canbe rapidly predicted. The procedures developed enable prediction ofpower-on longil@inal stalility characteristics that are generally invery good agreement.with experhnt.

INTRODUCTION

The prediction of thelongitudinal stabiliti andtractor airplemes has leen

effects of propeller operation on the static control ckacteristics of single-enginethe object of mmy investigations. Successfulmethods have been developed for estimatm the dtrect propeller forces

and the effects of slipstream on the wing-fuselage characteristics(references 1 to 4). Attem@ts to predict the complex changes h flowat the tail plane, however, have been 9omewhat less successful, primarilybecause many of the early researchers were hindered by insufficientexperimental data for developing methods of proven general applicability.

During the wsr years sn appreciable amount of experimental datapertaining to power effects on static longitudinal stabili~ was obtained.An analysis of these data suggested the possibilities of a semlempiricslapproach to the problem of detezmd.ningthe effects of power on the tailcontribution to stability. This approach has been followed in thepresent paper and a simple, rapid methd for determining the effects ofpower on the tail contribution is ~resented. Use of the proceduresdeveloped per?nitthe accurate prediction of power-on longitudinalstability and trim characteristics. No analysis has been tie for theflap-deflectedcondition.

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2 NW TN IJO. 1722

%.

Tc

TcA

ATc

PssbbDR

Ct7

lift

s-mBoIscoefficient

@tch~n* coefficientaverage section pitchhg+mm6nt coefficient aboutaerodynamic center fo~ whg section -rsed inslipstreamthrust coefficient. (m)thrust coefficient corresponding to powemff liftcoefficientincrement of thru8t coefficient from powe~ff condition

to a specified power conditionairspeed, feet per second .air denefm, slugs per culic footpropeller disk erea, square feetarea of wing or tail, square feetspan of * or tail, feetpropelle=lade section chofi, feetpropeilerpropellerradius towing meanwTng root

diameter, feetraditi, feetarw propeller %lade element, feetaerodynamic chord, feetchord at plane of symmetry, feet

w3ng chord at break fortigs having composite planforme, feetwing chord at theoretical tip, feetwing chord at spemwise station 0.50R or o.7X from airpl~e,centerline, feet

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NACA ~ NO. 1722 3

L ~ taper ratio (~/Cr for wings having ltiem taper)z distance fiam reference center of gravi~ to thrust ltie

meaeured perpendicular to thruet line (yositivewhen e.g. is almve thrust line), feet

ut

80

%0%4

S.F.F.

f

f.

distance frcm reference center of gravim to propellercenter line mea8ured yemllel to thrust line, feetdistance from reference center of gravi_&y*O elevator

hinge line, feetdistance fkam elevator hinge line to thrust ltne

me~ed perpendictilarto *t line (positivewhen elevator hinge line is above thrust line), feetangle of attack, radians unless otherwise denoted ..propeller blade angle, degreesstabilizer aetthg tith respectto thruet ltie (positivewhen trailing edge is down), degreeselevator setting with respect to chord line of staMlizer(positivewhen ~ing edge is down), degreeseffective angle of downwash at horizontal tail, degreesincrement of pwe-ff downwaeh at horizontal tail.fromzero lift dawnwash, degreespower+ff dowuuaeh angle at zero liftderivative.ofpropeller nomal-force coefficient tith

reepect to anglerfiane

valueof Cyt for+

of inclination of *tlline in

TC=O

abbreviation for propeller side-force factor

ratio of C-p$ to Oyt.*Oratio of ~t for power-off value of Tc to cytt*. o

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k IT4CATN No 1722

FRt

Re

ASubscripts:T

I etPPovwei

empirical tape=atio factord%ratio of power-on stabilizer effectiveness () P

to powe~ff stabilizer effectiveness ()dcmq.

()~atio of power-on elevator effectiveness d~eJ Pto power+ff elevator

change in a cpanti~ due

thrult 13neelevatorhorizontal tailyropellerpuuer onpower off

-tiel~ ccmibinatimimmersed fi slipstream

BASIS OF ANAmsIs

()l&effectivenesss dae oto power

The methcxiof coquting power-on pitcl+ng moments which isoutlined herein is lmsed on the assumption that power-off (propeller-off orwfnddlling) pitc~ t and lift data are available for atleast two stabilizer settings andtith the tail off. The accuracywith which the effects of power on the tail contribution to sta%ilimcan be predicted is dependent on the basic poww-off data, and whenpossible these data should be ohtalaed from w5@--tunnel tests.

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NACA ~ NO. 1722 5

When powe~ff w3nd-tUnnel data are not available for use in preliminarydesign, the pwer+ff characteristicsmay be estimated %y use of references 5to I-1. The w3ng mean aerodynamic chord and aerodynamic center may be foundby the methmi presented in reference 5. The liftiurve slope, angle of zerolift, and pitchin&mmn t characteristics of the wingmey he camputed ~y useof references 6 anii 7. The effect of the fuselage on the win@%selagepitching moments may Be detemined by Multhoppis method {references 8and 9). A satisfactory ap~oximation of the horimntal-tail effectivenesscan le obtained when the isolated horizontal-tail effectiveness found by themethod of reference 10 is multiplied hy a factor of 0.9.

The variation of power-off downwash with angle of attack computedbyuse of the charts of reference H generally was found to agree fairly wellwith the variation of effective downwash with angle of attack obtainedfrom wind+unnel data when the corquted downwash was multiplied by afactor of 0.9 for all conditions for averaging downwash across the tailspan instead of the factors Obt-a from figure 21 of reference I-1. Theabsolute angle of downwash computed by use of the charts of reference 11had to be adjusted, of course, so that this angle would agree with thetest downwash angle at zero lift. This adjustment was necess~ sincean appreciable amount of effective downwash was found to exist at zerolift chiefly as a &lrect result of local flow angularity at the tail.caused by the flow pattern over the rear of the fuselage. This downwash isoften difficult to predict accurately. Neglecting the downwash at zerolift, however, will not affect the basic longitudinal stabiliw or theestimated power effectf3but till alter only the tr- characteristicso

The exper-ntal data upon which the results of this paper are basedwere obtained fram wind-tugnel investigations of yowered models of specificmilitery airplanes. In figure 1 two views of each mdel and in table Ithe geometric ckacteristics of the configurations uEed are presented.Meet of the models were tested in the Langley yby &foot tunnel at an2effective Reynolds nmler of approximately 1.6 x 10 . The power-off datawere obtatied with the propeller win&WXnn at a value of Tc %+.015.Models 25 to 27 were tested in the Ames 7-by lo-foot tunnel at aneffective Reyaolds nmnber of approrlmately 2.0 x X36. The basic power-off data used for these ~~er mode~ ~re oh-d with the propellerlwnoved.

METHOD OF ANAIXXE

ti the following discussion the individual component effectscontributing to the ove~ power-on static longitudinal stabili~ aretreated separately and apprcdmate formulas are developed for estimatingthese effects.

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6

Effect of Power on the Win@?uselageDirect propeller effects. The incrament

.NACA TN

Characteristics

No. 1722

of lift coefficientcontributed %y the dtrect ~ropeller thrust due to the inclination of thethrust line (@e Mft cmponent of the normal force which isusually smallbeing neglected) is given by the following equation:

2 13inq%=Tc~ (1)The increment of pitching+moment coefficient contributed by thepropeller as a direct result of the thrust and the normal forces is givenby the following equation, which was developed from equation (5) of

reference 1:

(2)

where ~ is the absolute angle of attack (radians) of wing frcnnzerolift. Figqres needed for use in equation (2) have Men reproduced fromreference 1 and are presented as figmes 2 to ~. The term f - fo,. obtained from figure 2, is the difference in ~t +/%0 for power-onana power-off conditions. It shouldbe noted, however, that f. = Owhen powe~ff data are obtained with the propeller removed. Theterm Oyi@o o%tained from figur86 3 and 4, is the normal-forcederivative; figures 3(a) and k(a) ere for l~eed propellers having thick, cambered hlades; figures 3(b) and h(b) ere for high-speedpropellers having th@ ti,deblades; plan+omn curveg of propellers onwhich figwes 3 and 4 are based may be found in figure 4 of reference 1.The term de/da, the upwash factor, is obtained from figme 5.

Slipstream effect on win@hwelage characteristics.-The methodmost widely used for computing ihe increase h wing lift due to theslipstream is given by Smelt and Davies ti reference 3. lhiS methodreqties several successive approximations, however, to obtain finalpower-on lift coefficient when Tc veries with ~. An approximateformula has been developed which is shorter than that of reference 3and which requires only a single esthation to obtain the finaldm f %; thus an appreciable amount of computing time is saved.This equation is @ven ly

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ItfmA m Iio . 1722

where ~ is the wing chord at sp&ise station 0.75R frcm airplane centelhe for wings behhl single rotating propellers or 0.50R for wings behinddual rotating propellers.Very close agreament is shown between the values givenw equation (3)

and two approximations computed by the method of reference 3. (See fig. 6.Somewhat less agreement is shown between the values givenby equation (3)and test data (fig. 7). The scatter shown can he attributed to theidealized assumptions h the theoq of reference 3 and the experhntalinaccuracy of the test data. me effect of propeller tilt on AC~ issmaU as shown by reference J2, the data of model 24, and other unpublisheddata and may le neglected for tilt angles q to at least 5.

The effect of the slipstream on wing pitcMngmmnents is small insome cases, but it mey %e relatively large in others. This pitch~ntincrement isfollows:

where

/%q

[( )] lw~o

obtained tram equation (5) Cf reference 2 and is given as

[ (a)ti]c%oaverage section pitchti~nt coefficient aloutaerodynamic center for part of wing immersed in

slipstreamspan of fig imersed in slipstream (taken as 0.9D)average chord of wing immersed ti slipstream

rate of change of wing-fuselage pitching+mme ntcoefficient with lift coefficient (propeller off)

(4)

. . . . . . . .


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8 mcA m No 1722.Effect of I%wer an the TailContribution to Stablli@

-emduwnwashangl e due to power. The downwash atthe tail@Lane with the yroyeller removed is known to be chiefly dependant uponthe wing lift coefficient and the location of the tail with respect totie wing vortex system. When a propeller is added in front of the wing,q cqlex chsnges 3R flow occur which tiect tie downwash over thetail; lut the chief effects ere Pro%ably caused %y the increase hX@ lift coefficient and altered wing spa load distiilution tmoughtalout %y the passage of the yropeller 91iyt3&eem over the wing. Althoughappreciable a~ my etist %el&d en isolated propeller-at anangle of attack, lkrge changes in the inclination of the thrust -(at constant ~ angle of attack)-were found (reYerence13) to causepractica2Jy no change in effective downwash at the tail when a wingwas located %etween the propeller and the tail. With the foregoingdiscumion as a lasis the following simplified semieqiricsl approachwas used to derive a pmmeter tith which to correlate eqerimental.d~ changes au8 tO pwer.

Dounwash angles were ccunptiedfor several.models for whichextensive constmt thust data were availalle. When Ae was plotted.et Tc at various angles of attack, A~ was found to le afunction of loth Tc end a. This relationdliy f3eemd.logical forthe increase in wing lift with power, and any resultant Spen=loaa-es would slso deyend on Tc end aj huwever, c was believedto %e a better factor than a for use in the correlation Inasmuchas e usually varies linearly with a up to fairly high liftcoefficients end alao depends on tail location. -The assumption wasmade that a tdl well out of the powemff maximum a~ h fieldwOtia slso le favorally located in the yowez-on a0~8h field forconfigurationswithin the range of geq~ of the hmdels presented.

Model 17, which has an uatapered wing, showed a co-miderably largerincrease in e with the application of power than either model 13 ormodel 16, which hailidentical tail and fuselage geome~ but ratherhighly tapered wing plan forms. The power-off a~~ engles wereconsiderably less for the model with the untapered ylan form, but thea~ for All three wing plan forms cOtia Ye accurately c~utedfrom the charts of reference 11. With pwer, however, the.downwashangles for models 13, 16, ma 17 were much closer to the same value.

According to lifting-line theory the dm?nwashbehind awingofar%itrsry plan form h a uniform air stream (at a given tail locationand lift coefficient) depends only on the span load distii%ution alongthe wing. Wing taper ratio has a significant hearing on the span loaddistribution and hence the a~ h at the tail, the downwash increasingwith wing taper. The data of reference 14 show that the slipstreamslters the span load distribution ly increasing the loading over theinboard part of the wing. Since the greatest change fi downwash withtaper.occurs for values of X near un.i~, a wing of rectangular plan form

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9ACA TN NO 1722might be expected to be nore susceptible to this induced taper effectand show th e largest ohange of duwnwash angle with power. The experi-mental results of model-s13 and 17 qee with tie foregoing assumptions;however, further substantiationwuld be desirable.

An empirical factor F fig. 8) which is a function of w ng taperratio was derived frm the data of models 13, 16, 17, and other modelswith similar t~l geamptq to account for induced taper effects. Thetaper ratio for wings of composite plan formmsy be satisfactorilyestimated by use of an equivalenteroot chofi ~:s as.shown in the sketchti figure 8.

A plot of the paremeter (ATc)e*F against the exlermntal ~eobtained from stabilizer ad tail+f ~-tucnel data for 28 modelconfigurations is shuwn b figure g(a). The dash-line cmves h thefigure indicate the apprcmimate accuracy of determining downwash,angle~?&q complete+nodelwind-tunnel data. The correlation of test points Micates that the parameters selected account for the first+rdereffects of power rather well. The solidline curve indioates thesuggested cme for use in design and is reproduced in figure g(b) @th-out experimental test tit-a. ,

Note that figure g(b) inticates no &ange in A6 attributable to thetilt of the yropeller @rust ads. The tits of model 24 and other unpub-lished data show that ohanges .ti A G with tilt are small and ratherinconsistent and the effect of tilt (at least for tilt angles up to 50)oen be satisfactorilyestimated f~ considerations of direct thrusteffects ~ changes in stabilizer effectiveness.

change h stabilizer effectiveness tith power. The slipstream isconsiderably distorted in the region of the horizontal tail, andidealized theoretical methods which assume a cylindrical slipstream atthe tail do not always produce a satisfactory estimate of the changeIn dynemic pressure.at the tail associated with the application of power.As was true for the downwash correlation, a semiampirical approach wasfollowed to derive a method for predict- the change in stabilizereffectivenesswith power.

The ratio of power-on to power-off stabilizer effectiveness twas assmned to be directly proportional to ATc and the ratio of thepropeller diameter to tail span. A msxhum value of Rt was also assumedto he attained for the tail located on the thrust line with a lineardecline in Rt occurring until a value of unia is reached for a taillocation 1 propeller diameter ahve or below the &rust line. Experimentalpoints were plotted agairmt the paramecers s~sted by the foregoingassumptions (fig. 10) and the following relationship was obtained:

5)

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10 NACA T t?O. 1722

The dash-13ne cmwes ti figure 10 indicate the approximate limits ofaccuraq of detemdn.ing Rt from wind-tunnel test data, and essentiallyall the test yoints fdl within these limits.

Change in elevator effactiveness wfth power.- Tnasmuch as mostairplanes utilize the elevator for longitudhal control, it is apparent

~ tifi ~merthat the increase ti elevator effectiveness da will tnfluencethe determination of the final powemn stabili% and trim characteristicsof the airplene. An analysis was made to determine the possibili~ ofcorrelating Re h a manner similar to the foregotig correlation of Rt.The results of this analysis were less consistent than the results obtainedwith the correlation of Rt~ probably for the moat part because ofappreciable scale effects on some of the model elevator data and tiereduced accuracy possi%le in setting the elevatotieflection angles. Forfull~an elevators when estimated powemn elevator data are desired,however, it may le assumed in most instances that Re = ~.

Computation of Yowemn Lift and l?itchin@dament CoefficientsPower-on @n&%selage lift coefficient.- The final powe~n wing

fuselage lift coeffici~t is given by

w+=( )0+(AcbJ2% (6)b order to arrive at a value of

()?f pfrom equation-(6), the

followtng procedure is recommended for con.dit~onswhere Tc .varieswith ~: The tiCremant of the wing lift coefficient due to power isfirst evaluatedhy equation (3). The second-appro

T)tion of the increment

of W% coefficient contributed hy the propeller m? p is obtained bycomputing a first approximation ()ACLP by equation (1) with veluesof Tc corresponiMng to (Q). + Ah the second approximation isthen obtained by use of equation (1) with values of Tc correspondingh f%f)o + %, + (A%),. The second approximation will usually

(+)ive a value of C - such that further approximation will bePunnecessary.

. . - ..-. - --- - ,-. ~. - ,-., :- ,.


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mm m No 1722 u.

Tower-on wiog-fuselage pitching-moment coefficients. The finalpowe- wing-fuselage pitcMngaoment coefficient is given by

( ), = ( f)o % % (7)The terms A% and ~ are found from equations (2) and (4), withvalues of Tc heed on (%) as given by equation (6).P

Power-on tail pitc~ nt coefficient.- The computation of thepower-on tail pitching+mmnt coefficient merely consists of adding theincrements produced %y the altered dowmwash at the tail andincreasedtail effectiveness to (%) j this coefficient is given ~o

( ) = % ( -JO [()]%-AG Rtq 8)P oPowemn carplet~el pitchin~ nt and lift coefficients.

The fti power-on camplete+nodel pitcMnga&en t coefficient is givenby tiing equation (7) sm.dequation 8) as fofiows:

~, ( f)p+ (%)p

(%) ( ). .+P /% %the final Iowe~ complet~del lift coefficient is given byadding equation (6) and equation (10) as follows:

\ = (I?f)p-+w,

9)

(lo)

(lJ-)

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12

IllustrativeExample

mm TN No 1722

A detailed ste~yatep procedure for computing powe=n liftcoefficients and pitc~ment coefficients for model 21 is presentedin table II. The sample calculations h tabie ~ illustrate the useof the equations presented in this paper end give the pertinentconstants and column headings ti a convenient form for general applicationto deOign. The data from which these estimations were made and the finalcomputed poweron characteristics are prqsented in figures Il.and 12.Model 21 was chosen as an example because, although the indl.vidualcomponent effects of power were not small, the design v=iables weresuch that the adverse effects were counteracted.by the favorable effectsand thus a very small over-all change h stalili@ with power resulted.Calculations for models 13 and 15 were also made to show that the changeof power effects attilutable to raising the horizontal tail andincreasing its area can he accuratelypredicted. The basic data andestimations of power effects for these models - presented in fi~es 11,13, d 14.

The cmputed results for all three models show very god agreementwith the test results.

DISCUSSION

The range of the most pertinent geometric variables for the Incdelsused in this paper are presented in table ~. The correlation curvesof figures 6, 9, and 10 are bel-ievedto Be valid at least tetwe.entheMmits given in ta%le ICI. No data on powered models with appreciablewing sweep.were availa%lej consequently, the effect of sweep could notle included in the correlating parameters.

Wind.-tumneldata on permnal-@ye airplanes were not available foruse in the correlations, and the applica%ili@ of the curves Infigures 9 and 10 to this @e of design is de~endent on a nuuiberoffactors. Although the models used in the present comelat ion representhQhly powered fighter-@ye airplanes, the correlation curves werefound to be valid for medium power conditions on the fighte~eairplane mcilelsalso. An estdmate of the variation of Tc with ~for several @picel single-eme personal-e airplanes showed that thethrust coefficients for maximum rated qower for these airplanes fe12.inthe range of thrust coefficients for the medium power conditions on thefighte~e airplanes.

The range of wing vertical positions relative to the slipstreamand the ratio of the slipstream diameter to the wing span might beexpected to be c~iderably different for-militeryand personal-typeairplanes, and these differences could tive a significant bearing onthe magnitude of the power effects.

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/NACA TNNo 1722 u

KU. models presented.heretihave a wing looation that @aced thewing well within the slipstream. The increment of wing lift due to theslipstream derived from the data of reference 15 does not varyappreciably with wing height for wing positions 0.3 propeller diameterabove and below the thrust line when the propeller is more than 0.3 rootchord ahead of the wing leadtig edge. The range of wing vertical positionsfor the models presented herein is 0.165 and 0.176 propeller diametersabove and below the thrust line, respectively, and the propellers memore than 0.3 root chord ahead of the wing leading edgej thus the rangefallb within the l~ts of wing and propeller locations where coqputed

%vslues of A would be expected to.be valid.

Genera12y, the diameter of the propeller relative to the wing span issmaller for personalae airplanes than for fighter-@pe airplanes.Model 24 has a relative propeller diameter approximately the sane as forthe personal-e airplanes considered, but the other models used in thecorrelations had larger relative propeller diameters. No definiteconclusions can be made regarding the effect of relative propeller sizebecause of insufficient data. Inmost instances, the methods outlinedin the present paper should be satisfactory for computing the firstorder effects of propeller operation on personal-&ype single-enginetractor airplanes.

OPTIMUMDESIGN COI?SIDERMEHINS

The design configuration usually considered optimum when satisfactoryhandling qualities of airplanes are considered is that which exhibits nochange in longitudinal sta%ilim characteristics upon the application ofpower. Many design parameters affect the longitudinal stabili~ bothadversely and favorably, and defu a definite method by which to designan airplane with no power effects is difficult. Often considerations

t other than aerodynamic determine the final geom~ of a design. Zn viewof this fact and the rapidity with which the power effects of a specificconfiguration can be computed by the method of the present paper, eachproposed design should be examined for power effects, and an optimmconfiguration (mtnimnm power effects) should be attained by a process ofrational modification to the original design.Langley Aeronautical Laboratoq

National Advisory Committee for AeronauticsLeagley Field, Ta., Ju3y 13, 1948

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14,

NACA TN No 1722

1. Ribner, Herbert S.: Notesto Stabili@. NACA ARR

REFERENCES

on the Propeller and Slipstream in RelationNo. L4112a, 1944.

2. Pass, H. R.: Wind-Tunnel Study of the Effects ofPrope12er Operationand Flap Deflection on the Pitching Moments and Elevator HingeMomenta of a Singl&hgine pursuit-e Airplane. NACA ARR, July 1942.

3. Smelt, R., and Davies,.H.: Estimation of Increase h Lift Due toSlipstream. R. & M. No. 1788,British A.R.C., 1937.4. Goett, Hsrry J., and Pass, H. R.: Effect of Propeller Operation on thePitching Momenta of Slngl~ Monoplanes. NACA ACR, Mey 1941.5. Diehl, Walter S.: The Mean Aerodynamic Chord and.the Aerod3~lc Center

of a Tapered Wing. NACA Rep. No. ~1, 1942.6. Abbott, Ira H., Von Doenhoff, .Albert., and Stivers, Louis S., Jr.:Sumnary of Atrfoil Data. NACA ACR No. L5C05, 1945.7. ndersonwings .8. hitit.hop~,

R_nd F.: Determination of the Characteristics of TaperedlUICARep. No. 572, 1936.Ho.. Aero-cs of tie Fuselage. NACA @fNQ. 1036, 1942. .

9. White, Maurice D.: Estimation of Stick~ixed Neutral Points ofAirplanes. NACA CB NO. LXO1, 1945.10. Bates, William R.: Collection and Analysis of Wind=uunel Data on*theCharacteristics of Isolated Tail Surfaces with and without End Plates.,NACA TN No. 1291, 1947.I-1.Silversteti, Abe, endl?Latzoff,S.: Desi@-Cherts for Predict@ Down-wash Angles and Weke Characteristics behhd Plain and Flapped Wings.NACA Rep. No. 648, ~939.12. Goett, Harry J., and DelW, Noel K.: Effect of Tilt of the PropellerAxis on the Longitudinal-Stabiliw Characteristics of Sin@&EngineAirpl~s . NACA Rep. No. 774, 194-4.13. Sl%per, J.: Effeet of Propeller Slipstream on Wing and Tail. NACA TMNo. 874, 1938.14. Robtion, Russell G., and Herrnstein, William H., Jr.: Win&Nacel.le-

tiopeller bterference for Wings of Various Spans. Force andPressur@istiibution Tests. NACA Rep. No. 569, 1936.15. Wood, Donald H.: Tests of NaceD_~opell.er Conibinationsin VariousPositions uith Reference to Wings. Pert I. Thick Wing N.A.C.A. Cowled Nacelle Tkactor I&opellero NACA Rep. No. 415, 1932.

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19/43


20/43

JIMA m No 1T22 19

. TABLE llcI.-RAI?GEOF GEOMETRIC CHARACTERIHCTCSOF MODEL$ INCIUDED IN CORRELATION

RangeGecmmtric parameterFrom TO

Wing aspect ratio(mOz:i720) d: 19)

Wing taper ratio 0.275 1.00(model 19) (model 17)

I?ropellerdiameter o.217 0.354wing Spen (model 24) (model 20)Heightoof tail abate thrust line 0.413 -0.042

Propeller diameter (model 15) (model 11)Tail span 0.523 0.322wing span (model 3) (model 19)

Tail length 2.008 4.09Mean aero&namic chord (model 1) (model 9)Height of thrust lhe above wing root chord 0.165 -0.176l?ropellerdiameter (model 22) (model 23)Distance of propeller ahead of wing root chord 0.490 0.906Root chord (model 13) (model 21)

..-

.

...~-- ..- . ~..-- .. . -. .. . . __.,


21/43

m NACA II?O. 1722

.

Figure 1.- Model configurationssed inthecorrelations.

. . . .. . . -. -.. ,..... . .


22/43

a

1 10 II 12

1 13 1 14

1 17 1 18,

[ ( 1 I ()~ 17F== * -Figure 1.- Continued.

.. . ...


23/43

22 mm TN No 1722.

Figure 1.- Concluded.

. . . . .. .. . . -_.._-,., , .- .


24/43


25/43


26/43

Iuu2Am~0. 1722

l.k

.6

43 63 .% lm ml ml

(a)

Side-force factor, S.F.F.

Hamilton Standard prqmller S155-6; S.F.F. = &X7.

63 m la 1.40 la )Side-force factor, S.F.F.

(b) NACA propeller 10-3362045; S.F.F. - 1S1.6.Figure4.- Ratioofside-forcederivativesasa functionofside-forcefactorfordesiredpropeller.

---- ..---- ..-_ .. . . . . -. .- _.. ._. _ _. ._ ----- ..


27/43

I

2.0

1.6

L,Q

1 dq-z.8

.4

0-1.2 -.8 -. 4 0 .4 .8 1.2 1.6 . 2.o

Distance Mdnd root quarter -choxd point, rcmt chords

Figure 5.- Value of 1 - }: on longitudinalxk ofellipticing for aspect ratios of 6, 9, and 12.

toin

.


28/43

NACA ~ NO. 1722.,

.28

.24

-m

.ti

.I&2

.08

.04

0

I /AC Cw D2 0.57TC CL -~~- -Aoww\

Model

13 v17 a18 Q

23 ~

27

0 .1 .2 .3 . .4 .5 .6Cw D2T CL =CA OCw w

F@uxe 6.- Variationoftheincrementofwing Uftattributableothepropeller.slipstreamas computed from two appro&nations by themethod o+ -~ ~zreference3 withtheparameter Tc CL r .A 0~~

. .-... - - .. ..- .... . .


29/43


30/43

mcll m No 1722

1.0

.8

.6

F.4

.2

0

For wingshavingcompositeplanformCt A. Cr

Area 1 = Area 2

T ~LCt Crl

L . \ < J

0 .2 .4 .6 .8 1.0Taper ratio,A

Figure8.- Variationofempiricaltaper-ratiofactorwith taperratio.

..... . ..._.__ ._,_ ___ _______ .. . . _ _


31/43

,,

/

s

6

7

4

10 u3 ,, . 12 0

3/ % /

Q18 ~

123 ~

o

0. .4 .8 1.Q 1.6 Q.0 fi 2.4 2.8 3 .e

W&F

Figure 9.- Variation oftheincrement of downw=h angledue to power with the parameter (ATC),F.

.


32/43

.

6

9

4

3

2

1

0

II Ill 11111 1111111I I I I I I I I I I I I I I I I I

o .4 .8 1.2 1.6 2.0 e .h Q .8 3.2(ATc)dIF

(b) Swmted deelgn curve from_ B(a),

Figure 9.- Concluded.


33/43

32 NACA ,~ NO. 1722

P .6

2-4

2.Q

2.0

L8%

1.6

1.4

1.2

1.0

Model1234567891011113141516171819i)21

%2424a252627

0 .1 .2 .3 .4 .5- .6 .7 .8(ATc) @ t )t

Figure 10.- )ld~lVariationofexperimental Rt withtheparameter (ATC)$ 1 -~ .

. . . ..___ . ... . ..-. ,..-. . .. -- ----- -- .,. .


34/43

1.0

I

1I,I

iII

iI.1

0

ModelI I

13,16

21

/ /

/ /

/ /

/

/

o .4 .8 1.2 1..6 2.0

Lift coeftlcient, CL

Figure 11. - Variation of cmstant power thrust coefflcienk with Ilft coefficient for models 13,15, and 21. ww


35/43

.

_ ..

34 NACA ~ NO. 1722

-. Q

-. 3

1.6

-6-J+ o .4 .8 1.2 1.6

Lift Coeffkknt, CL

(a) *-tunnel data with propeller windmillhg.

Figure 12.- Longitudinalcharacteristicsfmodel 21.

.. ._,~. ,,,


36/43

ITAC.Am No 1722

.1

-. 3

1.6

.\ 3 -Taflofr

ft (deg)

0.

\ L

~ *.

-W=L T F 6.0

Test Compoted~

ft(deg)Ao8 %.

TalLotf

T

o .4 .6 1.2 1.6 2.0LiftCoemdent, CL

(b) Compulson of compukd and experlmenti data for constant ywer condlflon.

Figure 12.- Continued.

35

.-. -. .-.. .. . . . -- - . ..


37/43

.NACA ~ NO. 1722

3.0

0

-lo

T==?Fl

o .8 1.2CL

.

10

0

.9.

(c) Comparison ofcomputed and experimentaleffectsofpower on theneutralpointand tail-offerodynamic-enterlocation.Figure 12. - Concluded.

._ z .. - - . . ,,


38/43

Iwll m Ivo 1722 37

,

16

-8

Tailoff4 . . .xl -@ I+(deg)

e * IA 1.0

m

-El 5.83

+(d%)

A 1.0 5.890 Tmoff

IT

-.4 0 .4 .8 1.2 1.6LiftCoeffkklt, CL

,(a) Wind-tunnal data with propeller wWhUUn&

F@ure 13.- Longitudinalcharacteristics ofmodel 13.

-. -+. ..-. . -. .. .-. _ ... -.--


39/43

38 mm m No 1722

.

(b) Comparison of computd M experimental data for constant wer codltlon.

Figure 13.- Continued.

.

- . . ,-. . _,-, . ,


40/43

NAG A I IVNO.1722 39

*o

-10

-23

0

TestComputed

.

T

.+ .8 1.2Liftcoefficient,L

10

0

-10

g

A

(c) comparisonofcomputedand experimentaleffectsfpower on theneutralpetitand tail-offerodynamiccenterlocation.l&ure 13.- Concluded.

___..-, .. ____ . ~. .. - -- .. . -.,,


41/43

k)

-. 2

3.6

naca tn-1722 power effects

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