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TECHNICAL MEIvIOF&NDUIiS

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

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)F l?ROPEI,Ll123

NO* 874,.1.

—...———— -

SLIPSTREAM ON

J. Stiiper

Luftfahrtforschungvol. 15, No. 4, April 6,

von R. tifinchenOiden30uig,

.—-——— ——

WIITG AND

1938

TAIL

und Berlin

WashingtonAugust 1938

1

-\./’”” 1,,,.-,.. . .. —--————

NATIONAL ADVISORY COMI,llTTEX FOR AERONAUTICS

———..— —-

TECHITICAL KEMOI?AI?DUM NO. 874————

EFFECT OF PROPELLER SLIPSTREAM OITWING AND TAIL*

The results of wind-tunnel tests for the determina-tion of the effect of a jet on the lift and downwash of awing are presented in this report. In the first part, aset without rotation and with constant velocity distribu-tion is considered - the jet %eing produced by a speciallydesigned fan. Three-component, pressure distribution, anddownwash measurements were made and the results comyaredwith existing theory. The effect of a propeller slipstreamwas investigated in the second part. In the two cases thejet axis coincided with the undisturbed wind direction. Inthe third part the effect of the inclination of the propel-ler axis to the wing chord was considered, the results be-ing obtained for a model r~ing with running propeller.

I. INTRODUCTION

Attempts that have hitherto been made at constructinga useful theo,ry of the Iongitufi.inal stability of an air-plane in powered flight, have all come up against the dif-ficulty involved in the fact that the effect of the pro-peller slipstream on wing and tail has not yet been suffi-ciently investigated. In the present paper a study is madeof the mutual interaction of propeller, wing, and tail -the fuselage effect for the present not being investigated.The order of the three elements considered, namely, pro-peller, wing, and tail thu~ corresponds to the arrangementof multi-engine airplanes mith side engines. The problemsto be solved are the two following:

i

‘1.! iML;JI a) The effect of the propeller slipstream on the wing,, lift distrilmtion;I,j>*

!{i b) The effect of the propeller slipstream on the veloc-.~t[’ ity and direction of the flow at the tail loca-

tion.[,:‘1

1f’ .-.______---- _____------,,.-_,,---------------------..._.-,.-.. _._-.._,.,...____---_ ~._--.-. ,_

*“Xinfluss d.es SchraubeP-strahls auf Fl~~gel und Leitwerk. “Luftfahrtforschung, vol. 15, no. 4, April 6, 1938,

IIIi

pp. 181-205.I

.,.. — ——— -.— — ——

2 N.A. C.A. Technical Memorandum No. 8’74

The variable parameters of chief importance are: theangle of attack, the setting of the propeller axis to thezero-lift direction of the wing, the angle %ctween thepropeller axis and the relative mind direction, and thepropeller V/nD and thrust coefficient. The vertical po-sition of the wing in tho jet was not varied in our tests,the propeller axis always being on a level with the wing.The side engines in present-day airplanes are mounted ex-clusively in this manner, and in individual types the dif-ferences in the vertical locations of the engines amountat most to a value of the order of the wing-section thick-ness. As has %een shown ly both theory and experiment(reference 24), a slight displacement of this kind pro-duces no effec~ on th~ lift ~elations.

TABLE I.--——.———-—____

Airplane

———————————————-

Do 17He 111Ju 86Lockheed 12Lockheed 14Boeing 247-DBurnelli UB-14Douglas DC 3Douglas DC 4

Ha 139

Martin 130

Martin 156

Sikorsky S-42

——-—————.————__—

_—________ .____—__

~~t \

-——_—______ .___———__().92

68:74

1.02.84.68.66.82

inside .60out side .72inside .76outside .76inside .30outside .28inside .32outside ,34inside .42outside .42

-——..__. ._____—_—————

.—__—————ET

.-_—____——_

0.52.34.27.46.45.43,41● 43.33,40

● 21.21.26 ~.25.31.32.24.24

—————————-,

—

.————-_-———2A—-b

.——————————

0.28.23.25.27.28.21.15.19.19.41.23.49.13,32.11.29.12.33

,—————————

Talle I gives dimension ratios of the most importantrepresentatives of modern multi-engine airplanes, the no-tation being indicated on figure 1. It is evident, from aninspection of the table, that in spite of the variety ofthe types of airplanes, the numerical values indicated varyonly within narrow limits. The effect of a jet on a wingfor the case where the outside air velocity is zero - thatis, the problem of a wing spanning a free jet - has beenextensively investigated both theoretically and experimen-

.

3T.A.C.A. Technical Memorandum No. 874 3

tally (references 3, 10, 18, 19, 21, 23, and 24). The casewhere the qxternalair is in motion was treated theoretic-ally by C. llerrari (reference 7) and. C. Konj.ng (reference14). l?errari neglected to take into account the loundarycondition that must be satisfied at the jet-boundary sur-face . He found that...thcirculationon of the, wing-was%.,not

? -----,----@.a.?sed .l.w...,thyr.o~e,llerlersli~sti”~e-aiii~-’so”-”%hatthe increase

‘ in l~f-+ in the jet w~s”>r””o~-~r~~n>l to the increase in the@ovelocity in the jet. Konlng determined th&._jet__effect.<or

the case where the jet-boundary condition was satisfied.On account of the great mathematical difficulties involved,however, he was forced to make the following simplifyingassumptions. The angle between the relative wind direc-tion and the jet axis was equal to zero; the jet was freefrom rotational components and had constant velocity dis-tribution over the cross section and along the jet axis.Furthermore , it was necessary to assume tb.at the addition-al velocity in the jet was STIIaI-1compared with the maintunnel velocity in order that the problem might le “lin-earized.” A comparison of the investigations of Ferrariand Koning gives the remarkable result, namely, that therelatively rough computation of Ferrari leads to the cor-rect value for the total increase in lift produced hy the———._propeller slipstream, “out that the added J.l-fi=,~&t+~_~U~io.n@YQr the SP8.Q..de.viat.qs..st.r.~ngly...x.ornr<t<~e~eac,tual dlstribu--tL,ii.o”n”;‘-‘“-“~hfi-Sis explained the

---- . .good agreement of Ferrari ts

computations with the refiults of measurements where thedistribution of the lift is not taken into account. Nei-ther theory is capable of giving any clear predictions asto the actual downwash relations.~

The simplified assumptions of the theory of Koninglimit its usefulness for practical application. On theother hand, since the great multiplicity of factors in-volved - nonuniformity of velocity distribution aild rota-tion in the jet, mixing region at the jet boundary, incli-nation of propeller and jet to the relative wind, effectof the friction layer, etc. - make exhaustive theoreticaltreatment very difficult, if not impossible, the solutionof the problem must first be sought along experimentallines. For the purpose of learning the effect of the in-dividual parameters and their mutual interaction, the fol-lowing. investigation prog.ram,was form-dated.:

/

/;’ 1. Effect on a wing of a jet without rotation with

1’ constant velocity distribution, the relative wind direc-,1 tion and the jet axis coinciding (the case treated by

Koning); three component measurements, determination of

4 N.A.C. A.” Technical Memorandum 3T0..874

the lift distribution by pressure -dictrilnztion measurements,. determination of the do~nwash at the tail location.

2. Effect on a wing of -propeller slipstream with rota-tion, the propeller axis and the relative wind directioncoinciding; three-component and pressure-distribution meas-urements, determination of the downwash in magnitude anddirection.

3. Interaction of wing and propeller. The angle be-tween the propeller axis and the zero lift direction ofthe wing to he varied between 0° and 15°. Three-componentmeasurements and yressure-distribute on measurements, deter-mination of the downwash relations, effect of the propell-er V/nD.

4. Investigations on a twin-engine airplane in flight;pressure-distribution measurements in propeller sl-ipstream,downwash and long~tudinal-stability measurements, with par-ticular account taken of jet effect.

5. Six-component and pressure-distribution measure-ments on a model of a twin-engine airplane with propellerrunning; comparison with flight-test measurements, deter-mination of effect of direction o.f proneller rotation on

the down-mash and stability relations; ~ffect of angle be-tween propeller axis and plane of symmetry of airplane.

6. Determination of the shielding effect of an air-plane propeller operating with negative thrust.

The first three points of the above program are treatedin the present report.

II. VIN3-TUNNZL CORRECTION

The data presented in this report are the uncorrectedwind-tunnel measurements, since for the purposes of thiswork the application of such. correction wa~ not considerednecessary. For the sake of completeness, the values ofthe wind-tunnel corrections are given here. The correc-tion ACL, which must le applied to the angle of attack inorder that the lift in the bounded. tunnel flow equal thatof the infinitely exteiidcd flow, may be expressed as fol-lows :

—

N.A. C.A. Technical Mcmorandurn No. 874 5

where F is the win,g area

F0, tunne,l cross- sectional area

6*, correction factor

For the case of the two model uings here considered:

p = 0.16 m2

so that with F. = 1.765 m2

F/F. = 0.0906

In the computations made along the tunnel axis thereis obtained for the value at the center of the wings, tak-ing into account the end disk effect at the wing location:

.a) 6* = 1.000

so that au = 0.0113 c~ “

b) For the mean value over the wing:

and Aa = 0.0il_5 Ca

At the distance of 2.5 t behind the leacling edge of thewing, there is o%tained:

S* = 1.720

and Aa = 0.03.’94Ca

A further computation showed that the latter value

did not change appreciably within the limits of variationof the vertical location of the neasuring instruments forthe determination of the downwash relations (vaile, sphere,

,,two-prong apparatus)., It may be notqd in this connectionthat”.in determining the drag, correction was always madefor the drag of the end disks and fairings in addition tothat of the suspension members.

!“

6 N.A. C.A. Technical Memorandum No. 874

III. EFFECT ON A WING Ol?A JET WITHOUT ROTATION ‘i71TH

CONSTANT VELOCITY DISTRIBUTION

Test Set-up and Procedure

The mfiasurements were made on a wing of symmetricalsection (Gottingen 409) having a span b = 80 cm, chord ?= 20 cm between two circular end. disks of diameter h =32 cm. Figure 2 shows the test set-up for oltaining thepolar curves. The wing is suspended on a three-componentlalance in the wind tunnel of the propeller-research labor-atory. In front of the wing is the fan for producing therotation-free jet. Figure 3 shows a cross section of theblower. The main rotational components of the jet pro-duced hy propeller a arc removed by the vanes b. Thehoneycomb d further straightens the jet while the con-stancy of the velocity distribution is o%tained by suita-ble choice of the mesh e. The tQroat f -with the exitdiameter 2R = 12 cm has a slight flare at the end in or-

der to oppose the jet contraction. Figures 4 and 5 showthe dynamic pressure distribution for the two operatingconditions under tihich the tests were conducted (q = dy-namic pressure of undistur-oed flow, q. = dynamic pressurein jet). The measurements were made in the plane bisectedby the jet at various dj.stances x from the plane of thethroat outlet. Particularly to “DC noted is the only verysliSht increase in the extent of the mixinS zone with in-creasing distance from the throat. The velocity of theundisturbed flow Vm for all the measurements amounted to

30 m/s (6’7 m.p.h.). The relative increase s in the ve-locity in the equation:

v = (1 + s) Vm

(V is the velocity in the jet) has the values, respective-ly, 0.18 and 0.36. The friction boundary layer of the fanenclosure produces a dynamic pressure drop in the transi-tion regio~ between jet and undisturbed flow. It is to heassumed that this cyliilder of slowed-down velocity acts asa ccrta,in shield against the interfereilce of the flow proc-esses within and outside the jet. The jet was-,$~und t< be;y_:_efrom rotation.

-—.-,--, . ._..J..

In all the measurements the wing was located at thecenter of the jet; the changes in the angle of ,tittack wereeffected ‘by rotating the wing shout the leading edge. Tw O

N.A. IC.A. Technical Memorandum No, 874 7

.>

series of measurements were matle at distances of 0.25chord and 0.5 chord het~een the exit plane of the throatand’ the ‘wing leading edge. The re”sult”s”obtained we~*e iden-tical within the limits of accuracy.. The other measurementswere then continued at a distance of 0.25 chord.

Test Results

Figure 6 sho~~ the lift as a function of the angle ofattack. The jet gives rise to an increase

dcain the lift

slope~a-’

which increase, however, is not proportional

to the increase in the velocity. The maximum lift in-creases and f?.o~ separation is delayed to higher angles ofattack. This phenomenon is to Be explained by the effecton the je% of ,the boundary friction layer of the wing, thejet acting to delay separation. ~~lereas in the case ofthe wing without the jet the flow separation starts at thewing center, it is found that with the jet acting on thewing, separation starts outside of the jet region. Thisfact is of importance. In the design of an airplane theplan form and tmistfng Of the wing are so determined thatill flight at high ang].es of attack separation starts atthe wing center in order to nrevent dangerouS banking ofthe wing. This computation is generally conducted withouttaking into account the propeller slipstream. NOW the ef-fect of the slipstream is to support the flow at the cen-ter and the wing is again exposed to the d.an.gerof wing-tip stalling. The condition corresponding to power-onflight at large angles of attack is met with not only intake-off and climb but also in blind flying and in landingof seaplanes at full pol~er on smooth water.

Figure 7 shows the polar and moment curves. The jethas no effect on the moment curve and the polar shows aconstant increment of the drag as a result of the jet.

In order to be a-olc to determine the effect of thejet on the lift distribution, 17 orifices were lored ineach of 17 measuring stations along the s-pan. Figure 8shows the test set-up for t’he pressure-distribution meas-urements. The pressures at each of the measuring stationswere photographically recor”dad ~,~iththe aid of’ the multi-ple manometer seen in the foreground of the figure. Thepressure distrihution~ for the different angles of attackare given on figures 9, 10, 11, aqd 12, the pressuresli]qo (where p is the static pressure at the station)

,,,,, . ., .. . . ....-.————

8 N.A. C.A. ‘Technical “Memorandum No..8’74

being plotted against the projections of the orifices” onthe. undisturbed wind direction as abscissas. .The area in-

. eluded ly the pressure-distribution curves (obtained hyplanimeter) is thus the lift contributed at each section.The figures above the curves are the values of y/R (whereY is the spanwise coordinate) and hence indicate the posi-tion of each ,section from the jet center. The lift dis-tributions for the different angles of attack are shownon figure 13. It may %e seen that there is an increase inthe lift in the region of the jet %ut that in addition,there is a large effect also on the portions of the winglying outside the jet. The strong decrease in the liftis due to the friction boundary layer, already referred to,of the fan enclosed.

Comparison between Theory and Experiment

The first attempt to determine mathematically the ef-fect of the propeller slipstream on the lift distributionof a wing was made %y C. Koning (reference 14).

It is convenient here to give a short account of thetheory. To determine the lift distribution of a wing inparallel flow with propeller slipstream, the flow is di-vided into the following different parts:

1.

2.

3.

4.

5.

Theary are:

IIundistur%ed flC)TV,” with the velocity Vm.

llpropel~e~ flOW~” the difference between flow 1and the flow which would. exist if the propel-ler were acting in the absence of the wing.

“Wing flow, ‘1the,change in flow caused 3% thewing in the parallel flow in the a%sencc ofthe propeller.

“Additional airfoil flow,” the flow produced 3Ythe change in circulation of the disturbanceflow, related directly to the c?~ange in cir-culation around the wing, caused by the actionof the propeller.

I!Additional flo~! 1!the flow ~~hich is still to beadded to flows 1, 2, 3, 4,in order that the%oundary conditions may he satisfied.

conditions’ which must he satisfied at the hound-

IT.A. C.A. Technical Memorandum No. 874 9

a) The pressure must have the same value on eachside of the boundary, since there are no exter-nal forces acting on the jet boundary.,..

h) The component of the velocity normal to the bound-ary is equal to zero since the flow is free from‘sources and is steady.

Both conditions are satisfied by the flow components1 and 2, but in general, not by the components 3 and 4.In order that these conditions may be satisfied, it is nec-essary to add the additional flow component 5, and thedifficulty of the problem is just in determining this com-ponent . For the case here considered of a wing of finitespan lying aft of the propeller and intersecting the jet,an exact solution cannot he given and it is necessary tobe satisfied with an approximation.

In order to be able to carry out the computation atall, it was necessary for Koning to make a number of as-sumptions tvhich referred essentially to flow component 2.Figure 14 shows the position of the wing relative to thepropeller (jet) and to the undisturbed wind directionfor the general case. In the table below the simplifica-tions assumed in the theory are compared with the actualconditions.

——._..______ .__________________________

Parameter

———_——..—___—__— ___________

Anglo ‘oetwccn propelleraxis and wind direction

Form of jet

Velocity distribution V=:a) over the cross

section

b) along the jet axis

Velocity increase inthe jet

Jet structure

——.—.,-——--—.—————_-———.-

Actual condition

char-ges with imgleOf attack = CL-K

determined by jetcontraction

variable, fallingoff at the “edgeand center

variable, in theplane of thepropeller disk~r+o . ~ Vxm.

Uy to s z 0.8

‘.vithrotation

Simplificationby Koningtstheory———— _________ ..

zero

cylindrical

constant;v= (1+s) Vm

constant

s<< 1, sothat terms. S2 may~~ neglected

without rota-tion

,, , ,,.,,.m , ,.,, .!!-..,,, , u ,, , .,, ,.l, ,,, ,,,.-...-,, ,.., ,,, -.,! ,,.,, ,,, ,,,,,, ,,,-.,,. . . , , , . ., .--,,- ,,, ,,, . . ..- . . ——

I

. . . ,.

10 N.A. C.A. Technical Memorandum No. 874

Th’e.theory also takes no”account of the effect of the ‘jet on the processes in the frictional boundary layer ofthe wing. It is evident that these restricting assump-tions strongly limit the possibilities of practical appli-cation of the theory..

The case considered by Koning cannot he exactly repro-duced in any experiment. A strict proof of the theorywith the aid of the measurements conducted is thereforenot to be expected. The results nevertheless yield someinteresting data.

The continuous lines in figures 15 and 16 show thelift distribution as computed by Koning~s theory for thewing and jet used in the test. For the wing without jet(dotted curve) the lift distrilmtion was computed by themethod of I. Lotz. In the region of the jet %he measuredlift coefficients are smaller than “is “required %y the theo-ry. To a large extent this deviation is due to the 10SSin dynamic pressure due to the friction layer of the blowerbody as isalso shown by the sharp drop in lift at the jet%oundary. Outside of the jet the test results as well asthe theory show the surprisingly large effect of the jet.According to the theory, the relative increase in lift due

to the jetAca(y)-—----%-l.(Y)

is independent of the angle of attack,

the variation of this value along the span .for “our case be-ing shown in figure 17. If these values are compared with

the values ofAca(y)———— — o%tained from the measurements (fig*cao(Y)

18), the iatter show considerable dependence on the angleof attack. For this dependence on the angle of attack, tworeasons may he given. First, the effect of the, loss indynami~ pressure at the jet boundary depends very stronglyon the angle of attack as may be seen from fi.gizres15 and16. Secondly, the jet - whose diameter is smaller thanthe wing~ R/t = 0.6) -’‘is deformed with increasingangle of attack, the jet expanding on the pressure sideand contracting OD the suction side. The local lift in-crease %y the- jet thus %ecomGs smaller. It was possibleto verify the correctness of this supposition by testswith a jet of water. By means of air bu%bles the jet wasrendered visible, and thus the jet deformation caused %ya tving of varying chord could be observed. The betteragreement let~een theory and experiment at the larger an-gles of attack is also apparent,. The flow about the wingwithout jet %egins to break away, whereas in the presence

i.

——_ ..—.-.,,, —— ,m—mmmm nmmmmn I III - m I

N.A .C’.A. Tech?iiczLMemorandum’ No. “874’ 11

of the jet, separation occurs ‘later, and *his, conditionllca(~)

gives rise to ‘greater value’s”of:;:G> “9

TO the simplifying assumptions of Koning, there isthus.to be added a still further restriction, namely, thatthe ratio of wing chord to jet radius must be small enoughto maintain constancy of the jet cross “section.

Downwash Measurements

In order to obtain information on the flow”’relationsat the location of the tail do~n~ash measurements weremade in a plane at a distance of 2.5 chords behind thewing leading edge. Since the case of a jet without rota=tional velocity components with constant velocity distribu-tion is of less practical importance, we considered itsufficient in this case to determine the downwash anglewith the aid of a “feeler vanell only. This vane consistedof a metal plate 30 cm long by 10 cm deep, suspended on ascale behind the wing. Figure 19 shows the arrangementfor the downwash measurements behind the wing without fan;the vane can also %e made out on figuro 2. The measurementswere mado at four different vertical positions behind thewing. Figure 20 shows the. arrangement and the definitionof the symbols. Tho setting of the vane was so adjustedthat its lift vanished. Figure 21 gives the results ofthe measurements. The jet causes an increase in the down-wtash angle.

IV. EFXECT OF THE PROPELLER SLIPSTREAM ON THE WING

Test Set-Up, Procedure, and Results

In the following in~e~tigations the jet from the fanis replaced by that from.a propeller. The test set-up isshown on figure 22. A small high-speed electric motor en-closed in a wooden fairing, drives the propeller. Thelatter, has a diameter 2R = 15 cm, and a pitch H/D = 0.4(fig. 23> The measurements were conducted ,_ata propelleradvance ratio ~ =-0,15. Larger values “of ~ ‘we’re no-tused. since the propeller Would give no increase in velo,ci-ty on account of the ~alce of the streamlined body whichit would first be necessary for the propeller to acceler-ate. Before the.start of the measurementsp roper, the

12 N.A. C.A. Technical’ Memorandum No. 874&

propeller slipstream was investigated in the a3sence of thewing. Figure 24 shows tile variation of the angle of rota-tion, and figure 25 shows the dynamic pressure at variousdistances from the plane of the propeller disk. The dy-namic pressure was measured with a Prandtl-type tube, andthe angle of rotation with a two-pronged instrument. Inall of these flow-angle measurements, the angle given isalways the inclination of the flow to the horizontal. Thepropeller axis was fixed in the direction of the tunnelair velocity, and the angle between the zero-lift direc-tion of the wing and the propeller axis was thus equal tothe angle of attack. The distance between the plane ofthe propeller disk and the wing leading edge was 0.375chord.

The results of the three-component measurements aregiven in figures 26 and 27. In this case, too, there isan increase in the lift slope dca/da as a result of thejet. The two lift lines intersect, however, at the valueof the coefficient Ca = 0.2 (fig. 26). At smaller an-gles of attack therefore, the propeller slipstream leadsto a decrease in the lift. This phenomenon has also beenobserved from the results of various flight and model testswith power on (references 15, 17, 20). A closer examina-tion of this effect will be made in connection with thestudy of the lift distribution. (See below. ) As in thecase of the jet free from rotational components, the momentcurves of the wing are unchanged, whereas the polar showsa constant increase in the drag. Also in the presence ofthe propeller slipstream, separation of the flow on at-taining large angles of attack first occurred outside theregion of the jet.

The lift distribution was obtained by means of pres-sure-distrilmtion measurements. The distribution curvesare “shown on figures 28, 29, 30, and 31; figure 32 showsthe lift coefficients obtained from these. In the pres-ence of the propeller slipstream, two factors are effectivein changing the lift of the wing, namely, the “increase inthe dynamic pressure in the jet and the change in the rel-ative wind direction due” to the rotation of the propellerslipstream. The effect of the increase in,the dynamicpressure is proportional to the lift and, hence, approxi-mately proportional to the angle of attack, while the ef-fect of the rotation is, in general, independent of theangle of attack as long as the linear portion of the liftcurve is being considered. As may le seen from figures24 and 25, the dynamic pressure increase and the jet rota-

hT.A.C.A. Technical Memorandum ITo. 874 13

tion act in the same sense (tending to increase the lift)on the right side (positive values of Y), while on thele”ft side’ the rotation an”~le .an”dtho increase in the dy-namic pressure oppose each other in their offcct on thewing. Whether the one or the other effect prevails, de-pends on the angle of attack as ma g le seen from figure 32.For angles of attack up to about 8 the, effect of the ro-tation angle is predominant - the propeller slipstreamproducing on the left portion a decrease in lift a.s com-pared with the wing in the absence of the propeller. Thiseffect may therefore be strong enough so that for thesmaller angles of attack the $otal change in lift may evenbecome negative as a result of the propeller slipstream.

Domnwash Measurements

With the arrangement indicated above for the feelervane (figs. 22 and 33), the d.otvnwashwas measured at thelocation of the tail. The results presented in figure 34show an unexpected decrease in the tiotvnwashangle duc tothe propeller slipstream. This result which, on repeatingthe” test, prove~ to be reproducible, stands in contradic-tion to practical experience and model tests (reference 5),which always give an increase in the dotvnwash angle. Theexplanation is -prolably to be found in the fact that whilethe total lift at the vane vanishes, the lift may not van-ish locally everywhere, The measuring vane is relativelylargo compared with the jet dimensions (vane span 30 cm,propeller diameter 15 cm), so that a consid.crable portionof the vane lies in the upwash near the jet. In tho caseof the model measurements referred to above, the ttail waslocated entirely in the prope].ler slipstream.

In order to study with sufficient accuracy the effectof the propeller slipstream on the flow at the tail, andalso to investigate the jet itself, a survey of the flowfield in magnitude and direction was made in a plane nor-mal to the wind” direction at 2* chords behind the leadingedge of the wing. Figure 35. shows the l-ocation of themeasuring plane. As a &o’ntrol and for applying a correc-

tion, there was first measured the flow direction from thetunnel alone, the tunnel. flow being found free from ro-tational components. There was then determined the down-wash angle behind the wing in the absence of the propellerat the”’two positions y~t = ~ 0.533 (fig. 36). The “val-leysll shown”on the curves as shifting upward wit,h increas-in~ angle of attack are due to the domnwash from the wing.

14 IIJ.A. C.A. Technical Memorandum No. 8’74

The downwash relations for the various angles of attackare given in figures 37”, 38,. 39, 40, and 41. The cross-hatched areas represent the change in the dynamic pres-sure ; they bring out the fact, particularly noteworthy,that the propeller slipstream is cut %y the wing into twoparts which do not again unite into a single jet. In ac-cordance with the jet rotatidn, the upper portion is devi-ated to the left and the lower portion to the right. Com-parison with figures 24 and 36 shows that the wing removosa considerable portion of the rotational motion in thepropeller slipstream and this acts to some extent as aflow straightener.

V. hfiUTUAL INTERACTION OF WING AND PROPELLER

Test Set-Up and Procedure

In the following tests a study was made of tho mutual

interaction of wins and propeller, the angle between thezero-lift direction and the propeller axis being variedhptween 0° and 15°. Yor this purpose it was necessary tomake an arrangement whereby the propollcr and wing couldinteract without any outside disturbance. An undesirableeffect would have been o%tained, for example, if the driv-ing motor for the propeller mere located in a nacelle atthe wing. The previously employed arrangement of enclos-ing the motor in a fairing ahead of the wing would, inthe present case, have led to difficulties in mounting andundesired effects on the flow since, with changes in angleof attack of the wing, the propeller axis would correspond-ingly have to %e rotated along. The inclined flow on themotor body would have given rise to considerable disturb-ance .

~i~ure 42 ~homs the model used in the t~st. The no-tor is attached outside of the flow to an end disk anddrives the propeller through a pair of bevel gears and,ashaft located in the wing. Figure 43 shows the wing with-out, and figure 44 with, motor,,enclosed in the fairing.The wing, of profile section Gottingen 398, has a spanb = 80 cm and a chord t = 20 cm, with end disk diameterh = 32 cm. Figure 45 shows the lift curve of the wingalone , and f~~re 46$ the polar. The coefficients arogiven in table II. To carry out the pressure-distributionmeasurements, 20 measuring stations wit,h 14 orifices ea~were distributed over the span. The propeller shaft was

IT.A. C.A. Technical Memorandum No. 874 15

located in a bearing piece (fig. 42). By interchangingth%s. leaq~ng v$th. others, it was possible to obtain dif-ferent settings K of the wing chord to the p“rOP”elleraxis. Four such pieces were used (fig. 47). On figure48 are indicated the two extreme positions of the propel-ler axis and the de~inition of K; the four values of Kused mere 9°, 4°, -1°, and -6°.

The propeller used in these tests is shown on figure49, and its thrust, torque, and efficiency curves aregiven on figure 50. With the arrangement employed, H/D =1.0 and 2R/t = 1.034. In the three-component measure-ments the values of A used were 0.13, 0~16, 0.20, 0.35,and 0.55; while in the pressure-distribution and downwashmeasurements the ve.lue of 0.13 tias omitted. The tunmlair velocity ,in all cases was 30 m/s.

TeS’t Results

‘The numerical values of the three-component ?neasure-ments are given in table III. Figures 51, 52, 53, and 54show the variation of the lift witlh angle of attack. Forthe purpose of discussion of the results, it is to benoted that the total lift measured on tile scale was madeup of four component parts:

1. The lift from the wing itself, AIY

2. The lift at the ~in,g due to the propeller slip-stream, Ast.

30 The component of the propeller thrust in the liftdirection, AT.

4. The lift due to t]l~ inclined tunnel flow on thepropeller, AL.

In general, the lift of the wing Am ly’ far exceedsthe other components. Of the other three components, animportant part with regard to the forces is played’ - ex-cept in extreme cases only - by” that due to the propellerslipstream (ASt), while the othertwo (AT and AL) mayhe neglected. Hovever, in the. study of the moment equi-librium about the lateral axis, the two forces ATAL

andare of significance since they generally act on a rel-

atively large lever arm., Whereas the lift of the wingdepends essentially on”the angle of attack,. in the case of

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

16 N.A. C.A. Technical Memorand.urn No. 874

the other three components there also enter s”thethe angle K since the vropcller is attacked at

effect ofthe an-

gle -a-l%. For the nor~al-range of values of a and Kthe following may be stated. With increasing value of K,

Ast becomes larger siilce the angle at which the wing is

attacked by the propeller slipstrce,m becomes larger; AT

tind AL become smaller, however, since the angle betweenthe propeller axis and the wind direction (= a - K) be-comes smaller.

Figures 51, 52, 53, and 54 show the increase in thedca

lift slope ~a– through the effect of the propeller. Tb.e

effect described above - namely, that .% smaller.angles ofattack the jet leads to a decrease in the lift - may also

he observed in this case. The position of the point ofintersection of the lift lines depends, however, on theangle K, and the decrease :.nlift hccomes less with de-creasing K. From the consideration on the lift distri-bution, it may be coilcluded that in the region of smallangles of attack for larger values of K, ASt > AT + AL.

Furthermore , it is to be cxpcctcd that at small values ofh ,that portion of the lift contributed directly by thepropeller (AT + AL) gains in importance and that with

increasing K, the total lift becomes smaller. Figure 55,showinE the lift curves at A = 0.13 for various valuesOf K, confirms this prediction.

&caFigure 56 shows ~~– as a function of K, and fig-

‘t ma~y beure 57 as a function of A. . seen that the ef-fect of the angle K on the total lift is not large.This fact comes out even more clearly when the polars arestudied (figs. 58, 59, 60, and 61). Figure 62 shows thepolars for A = 0.13, 0.16, and 0.20 for various valuesof K. With the exception of the polar forK 90

~ = 0.13 and= the curves almost all coincide. This means there-

fore that in varying the angle K within the prescribedlimits, the individual effects (slipstream, inclined pro-peller, etc.) vary, but the sum of the effects on the en-tire wing-propeller system remains constant.

I!xcept for the maximum lift region, in passing frOmone value of A. to another, the value of Cw changes by

an amount which is inde~endent of the lift coefficient(Ca); that is, for cha~ges in ~ the polars shift

the ‘w axis. Starting from the polar of the wingthe value Acw by which cm changes, is a measure

alongalone,of

N.A. C.A. Technical Memorandum No. 874 17

the propeller thrust. The disk loading of the propeller

sc~ = ——-

~s qo= + ~ LJ... z= ;$:~

(s = thrust ,

Fs = propeller disk area) z ~~ti T—.. .—+ ‘“ -@&’ ‘+,:’;”

is obtained from the relation:

FCs = A% (F =

F;wing area)

Figure 63 shows c~ as a f-unction of ~ for the propel-

ler alone and for the propeller. in the presence of thewing. The difference between the two curves gives theinterference effect of wing and propeller~

———————.. .—

a———__— _____

,-8°

-6°

-40

-2°

4°

6°

8°

10°

120

140

16°

18°

________________ .

Ca.——-———-..—— ... .... . . . .

--C.102

.G40

.196

.351

.502

,661

.816

.965

1.104

1,231

1.342

1.413

1.434

1.376

-————_———--————

%...——————-.——-——

0.012

*G07

.010

.012

.018

.028

.041

.060

.078

.099

● 119

.142

.168

.209

-—-———— -————

cm-—————— —.—————

0 ● 044

.089

.121

.164

.195

,240

.285

.394

.419

.452

,.460

.478

N.A. C.A. Technical Memorandum No. 874

The moment curves (with respect to the wing leadingedge) are shown in figures 64, 65, 66, and 67. The un-stalilizi.ng effect of the propeller may be seen first fromthe increas= in the values of cmo s and secondly, from

the lowering in dcm/dca with increasing values of h.

No effect of the angle K on the values of dcm/dca

could be made out. (See fig. 68. ) I?igure 69 shows thevariation of the stability coefficient d~m/ dca of the

wing with propeller. Figure ’70 shows the value dcm/dA,which plays an important part in the theory of longitude-’nal stability, as a function of A. In the determinationof these- values, no dependence on ca was found withinthe limits of accuracy employed.

For the determination of the lift distribution alongthe span, pressure-distribution measurements were carried out,the test set-up corresponding to the one already described.The measurements mere made “at the angles of attack whichcorrespond to the main flight conditions: high-speed flight(a = ‘3°) and take-off and climb (a = 8“). The pressure.distribution curves are in this case not given since theircharacter does not differ from the curves given in thepreceding sections. Figures 71, 72, ’73, and 74 show thespanwise lift distributions. For operating conditions, inwhich the propeller produces a thrust (A = 0.16 and 0.20),the propeller slipstream giyes rise to a strong increasein lift, whereas in the case where the propeller is oper-ating as a nindmill (h = 0.55), the propeller slipstreamresults in a lowering of the lift. To the left sides ofthe figures the dynamic pressure increase and the angle ofjet rotation act with opposite effect on the wing, andthis explains the llunrestll in the lift distributions, par-ticularly at the jet boundary (y/R ~ -1), where thevortices separating from the propeller-blade tips are lo-cated.

In the pressure-distri’oution measurements the valueof Am + ASt is measured as the lift. The effect of theangle K on the lift distribution must therefore 3e takeninto account sincm ASt depends on K, although Aw \

does not. This effec’t of K may be clearly made out onthe figures and is more evident in figure 75, which showsthe lift distribution for A = 0.16 ‘with different val-ues of K. With increasing values of K, the increase inlift as a result of the jet is greater since the direc-tion of the jet causes an increase in the effective angleat the wing center.

... - .-—.—,.. ,. m -..,, . . ,, .,, ,,, ,,, ,, ,,, ,,, , ,,, , ,,---- , ,,, , , ,,,---

N.A. C.A. Technical. itie.m~randum.No. 874 19

.,. Of particular ,importance is the separation process onattaining, large angles o.f attack.; In ,order to study thisprocess, tuft investigations,were made on the wing”witband without propeller, and photographs also obtained on afilm. Figure 76 presents the results of these tests. Atthe crosshatched areas the flow has separated. The lefthalf of the figure shows hom separation at the wing alonebegins at the trailing edge of the wing center and fromthere on spreads over the entire wing. In the case of thewing in the presence of the propel&er on the right half ofthe figure, the value of K was 4 and A = 0.16. Separa-tion starts at the trailing edge at the positions.of the jetboundary, and from there on the separation is propagatedtoward the wing tips, whereas in the jet region itself theflow continues to adhere far beyond the maximum lift. ITOeffect of the nonsymmetry a.ue to the propeller rotationcould clearly %e made out on the separation process.

Downwash Measurements

Iila plane 2+- chords behind the wing leading edgethe downwash was measured in direction and magnitude withthe aid of a dynamic pressure sphere. The test set-up isshown on figure 77, and fi~ure 78 shows the relative dimen-sions. The measurements were made along two horizontallines: one in the projection of the ming chord (positionI), the other 0029 chord shove the latter (position II).In changing the angle of attack of the wing the positionof the sphere was like~~ise all~ays changed to correspond tothe rigid arrangement of wing and tail.

The moment MH of a horizontal tail surface ‘is, withthe usual notation

‘H = ~o rl~ (~-~)%FHi c!

and hence the stability contribution of the tail

The factors in front “of the brackets are design values ofthe tail while the expression ~vithin the brackets is ameasure of the “quality of the flOW at the position of thetail. This value we shall” denote by E. It is immediate-ly evident that for an’~le~ator in a nondistu~bed flow

,,:.. ,.,,

.,

20 N.A. C.A. Technical Memorandum No. 8’74

c 1. The terms elservhere proposed of “tail efficiency”or=tlstability efficiency itdo not correctly bring out thesignificance of E ; a better term would appear to hellefficicncy of the tail flow.l[ Therefore, we have:

In our measurements, in which A and q. were held coi~-

stant with change in angle “of attack,

may approximately %e set equal to zero, so that

‘=(’-%);:Figure 79 shows the domnwash for the wing without

propeller. At the angle of attack a = 16°, the flotih~d already separated. The small downwash value at a =8 arises from the fact that in this case the sphere waslocated in the dead-air, region of the wifig. The effect ofthe propeller on the downvash relations is shown on fig-ures 80, 81, 82, and 83. The crosshatched areas give thechanges in the dynamic pressure. The values shown are fora=- 3° and 8°. In order to include the effect of theinclination of the propeller to the wing chord the meas-urements were taken for K = 9° and -6°.

In the study of the domn~ash, it is to be notqd thatseveral factors determine the flow behind the wing l~ithpropeller, namely, the down-wash of the wing itself, thedead-air region of the wing, the locally limited propellerslipstream with rotation and variation in” dynamic pres-sure, and the effect of the slipstream on the flow in itsneighborhood. According to the angle of attack the pro-peller’ slipstream will envelop the entire tail or only apartof it, or may pass above or below it. The dead.-airregion of the’wing leads in general to a decrease in thedo~n~ash and the dynamic pressure. The shape of the.dead-air region is changed by the jet. Pigures 80 to 83 showthe interaction of all these factors. In position I (inthe projection of the wing chord) the jet effect may he

NeAoc. Ao Technical Memorandum No. 874 21

made out in the case of all four measurements, whereas in~ositi.on .11,.particula.rly at a,= -3°, the direct jet ef-fect is vanishingly small. Z!lieupper half of figure 80brings out .-theeffect of the dead-air re~ion of the wing.l?igures 84 and 85 show the variation of E , the l’efficicn-cy of the tail flowlf al”on& the span. Difficulties were metwith in determining 38/Z3a since the value of 6 verymuch depends on Which of the ~lany factors mentioned aboveis predominant at the particular position. It is impossi-ble to make any definite statement as to whether the sta-b~.lity contrilmtion of a tail surface in the flow investi-gated is diminished %y tb.e effect of the propeller slip-stream. In all the meas~~.remenfjsit may clearly be madeout t~hat t~.ere exists upl~ash near the jet. ~The inclina-tion of the propell.e_r_-~_~_is(K) has no demonstrable Eli-.......f~n the .]own~ash, whiE17%E~Z~=---”-’–”-–-”

__——. -.._...—--------in good agreement with

the”-cofistancy found for the total lift. .

VI. SUMMARY

In the first part of the investigation the effect ona wing of a jet without rotation with constant velocitydistribution, is determined. The jet gives rise to an in-crease in the lift. i~o accurate check on the theory ofI{oning, which unfi.erli~~ this case, COUld be undertakensince some of the assumptions made in the theory cannot besatisfied in the test, The downwash measurements at thetail location showed an increase in the down~rash angle dueto the jet.

In the second Fart of the investigation the wing wasunder the effect of the jet from e,propeller whose axiswas fixed in the direction of the undisturbed wind. Therotation and the dynamic pressure c~anges in the jet re-sult in a nonsy~mmotrical variation in the lift. study ofthe downwash relations led to the result that the two por-tions into ~~hich the jet is divided Yy the ~Ting do not .again reunite %ehind the wing but that each -portion experi-ences a la’teral deviation in the direction of the jet ro-tation.

In the third part, the mutual interaction of wing andpropeller was investigated. The propeller shaft, whichwas driven by a motor attached outside the wing itself,could be inclined with respect to the wing chord-. Thisinclination has considerable effect on the change in liftof the wing by the propeller slipstream. The total lift of

/’

22 lT.A. C.A. Technical Memorandum No. 874

>

the wing-propeller system in which lift is included be~sides that. of the wing proper, the component of the pro-peller thru’st in the lift direction and the lift due tothe inclined position of the propeller with respect to thewind direction, is hardly affected by the inclination ofthe propeller-to the wing chord, and similarly, no effectcould be established on the moment curve. The propellerincreases the instability of the wing. By dovnwash meas-urements it was determined to what extent the characterof the flow at the tail is changed under the effect of the

b

propeller slipstream.

‘Translation by S. Reiss,National Advisory Committeefor Aeronautics.

1.

2.

3.

4.

5.

6.

7.

Blenk, H.: IUight Tests for the Determination ofStatic Longitudinal Stability. T.M. l~Oo 584,lT.A.C.A., 1930.

Blenk, H.: Luftschrauhenstrahl und L~ngsstabilit~t.Luftfahrtforschurig, vol. 11, 1935, p. 202.

Illenk, H., and Fuchs, D.: Druckmessungen an e,inemdurch einen Luftstrahl hiildurchgesteckten Trag-flugel. DVL Jahruch, 1931.

Bradfield, F. B.: Preliminary Tests on the Effect onthe Lift of aWing of the Position of the AirscrewsRelative to It. R. & M. No. 1212, British A.R.C.,1928.

Bradfield, F. B.: Wind Tunnel Data on the Effect ofSlipstream on the Down-wash and Velocity at thaTailplane. R. & M. No. 1488, British A.R.C., 1-932.

Ebert, H.: fiber I’lugversuche zur Messung der Flug~zeugpolare und den Einfluss des Schrau%enstrahls.DVL Jahrbuch, 1932.

Ferrari, C. : fiber den Einfluss der Luftsch~aube aufdie aerodynamischen IZigenschaften des 3’lugels.LIAerotecnica, vol. 13, 1933.

N.A. C.A. Technical Memorandum No. 874 23

8.

9.

10.

11.

12.

13.

14.

15.

16.

17.

18.

19.

20.

l?lachs%art, O., and Kr~%er, G.: Experimental Inves-tigation of Aircraft Propellers Exposed to Ob-lique Air Currents. T.M. No. 562, N.-4.C.A., 1930.

Gorski, V. P.: Untersuch~ng ~her den EinflussdesRumpfes und der Tragflache auf horizontal Schwanz-fl~chen d.es I?lugzeugs. CAHI Report 131, 1932.

Glauert, H.: The Lift and Drag of a Wing Spanning aYree Jet. R, & M. No. 1603, British A.R.C., 1934.

H~bne&, W.: ities~ungdel* H~hensteucrkr~fte und derL~ngssta.-Dilitat eines I’lugzeugs vom Muster, JunkersF 15 ge. DTL Jahrbuch, 1.930.

E{bncr, W.: Erge3nisse von Messungen der Stabilit~turn die Querachse. DVL Jahrhuch, 1931, p. 684.

H{3ner, W.: Erge%nisse von Mcssungen dcr statischenL&gsstabilit&t einiger Flugzeuge. DVL Jahrhuch,1933.

Koning, C.: Influence of the Propell.er on the OtherParts of the Airplane Structure. Durand, Aero-dynamic Theory, vol. 4, 1935.

Millikan, C. ~~, Russell, J. S. , and MCCOY, H. M. :Wind Tunnel Tests on a High Wing Honoplane. Jour.of the Aeroilautica~ Sciences, vol. 3, no. 3, 1936.

Misztal, T.: Zur Frage der schr~g ,angeblasenen Pro-peller. Abh. aus dem Aerod. Inst. Aachen, no. 11,1929.

Ostoslavsky, I., and, Halesoff, D.: Interference be-tween Airscrew and Aeroplane. CAHI Report 213,1935.

Pistolesi, E.: Sull!ala traversante un getto libero;Atti dells Pontificia Academia delle Scienze,NU.OVU Lincei , vol. 86, 1933.

Pistolesi, E.”: L1influsso dells limitazione deilecorrente sulle cratteristiche dei modolli di ali.LIAerotecnica, vol. 16, 1936.

Pleines, W.: Flugmessungen im H~chstauftriebsbere ichmit dem Flugzeug Focke-Wulf A 32 “Buzzard”.Luftfahrtforschung, yol. 12, July 30, 1935, p. 142.

1’ ..- —

24

21.

22.

23.

24.

25.

#

26.

N.A. C.A.

Prandtl, L.:handlungenL. Prandtl

Technical Memorandum No. 874

Tragfl#geltheorie . Neudruck, Vier Ab-zur Hydrodynami~ und Aerodynamic vonund A, .Betz. Gottingen, 1927.

Robinson, R. G., and Eerrnstein, W. H.: Wing-Nacelle-Propeller Interference for Wings of Various Spans.Force and Pressure-Distribution Tests. T.R. NO.569, N.A.C.A., 1936.

St{per, J.: An Airfoil Spanning an Open Jet. T.M.ITo. ‘723, N.A.C.A., 1933.

St{per, J.: Contribution ta the Problem of AirfoilsSpanning a Free Jet. T.M. No. ‘796, N.A.C.A., 1’336.

Wood D. H., McHu,gh, J. G., Valentine, E. 3’., andB~oletti, C.: Tests of Nacelle-Propeller Combina-tions in Various Positions with Reference to Wings.

Tra,ctor-Propeller, by D. H. Wood.T.R. No. 415, N.A.C.A:, 1932.

II - Thick Wing - Various Radial-Engine Cowl-ings - ‘Tractor Propeller, by D. H.Wood. T.Il. No. 436, i{.A.C.A. , 1932.

Part’ I - Thick Winz - N.A.C.A. Cowlod. Nacelle -

Part

Part

Part

Part

Part

III - Clark Y Wing - Various Radial-EngineCowlings - Tractor Propeller, by D.H. Wood. T.R. No. 462, N.A.C.A. , 1933.

IV - Thick Wing - Various Radial-Engine Cowl-ings - Tandem Propellers, by J. G.McHugh. T.R. NO. 505, N.A.C.A., 1934.

v; Clark Y Biplane Cellule - .N.A.C.A.Cowled Nacelle - Tractor Propeller,by E. ??. Valentine. T.R. No. 506,N.A.C.A., 1934.

VI - Wings and Na,celles with Pusher Propel-ler, by D. H. Wood and C. Bioletti.T.E. NO. 507, fl.A.C.A. , 1934.

Lotz, I.: Einfluss des Sch&aubenstrahls auf die Auf-trielsverteilung. Jahrbuch, 1936, der” Lil~enthal- ‘Gesellschaft. Verlag von R. 01$.en%ourg, Munchenund Berlin.

I1 K=90

Table3,

a

l_80

cd

— 0,2470.146

c,” c,,,

—0,441 0,101—0,450 0,183—0,443 0,251—0,418 0,325—0,364 0,399—0,266 0,483— 0,2240,531— 0,1430,535

— 0,Z47o,07i— 0,2510,167— 0,2410,233— 0,2130,316— 0,1630,387— 0,1040,465— 0,0450,5150,0030,522—

— 0,1290,067— 0,1340,149— 0,1200,225— 0,0650,302— 0,0490,3860,0100,4660,0580,5010,1100,511

c,, c0“ c,,, cd c,” c,,,

— 0,154— 0,4250,1320,244— 0,4300,203

c“ c,. cm

— 0,2110,1780,6610,9551,3161,6451,8481,815

— 0,418— 0,427— 0,411— 0,374— 0,324— 0,271— 0,183— 0,130

0,1040,1840,2620,3310,4220,4800,6450,676

— 0,1920,2100,5910,9681,3351,6721,8601,820

— 0,420— 0,433— 0,421— 0,376— 0,319— 0,259— 0,165— 0,107

0,1230,1990,262C$3450,4250,4850,5450,586

— 4000.408°12?16018°

— 8°— 40004fi8°1216°180

_ 80_400040801%16°18°

_ 80_ 400040801916°18°

_ 80_ 40

(-JO4080

1P16°18°

0;5300,9061,2931,6211,8311,803

— 0,1950,174.0,5260,8991,2361,5361,7151,651

– 0,1650,1830,5220,8621,1841,4871,6281,562

0;639— 0;4160;2701.030— 0.3720,352a=o,131;3761,6961,8801,843

— 0;3060;432— 0,2210,502— 0,1380,560— 0,0700,592

— 0,1670,1910,5420,9041,2361,5361,7081.664

—.—— 0,249— 0,254— 0,237— 0,212— 0,170— 0,106— 0,0350,008

—.— 0,133.0,236

— 0,2580,093— 0,2580,171— 0,2410,230

0,0880,1550,2340,3110,3860,4490,4900,529

0,0720,1490,2200,2990,369

—0,1480,2020,5540,9221,2701,6611,7111,660

— 0,257— 0,258— 0,240— 0,206— 0,166— 0,101— 0,0220,023

— 0;135— 0,135— 0,120— 0,092— 0,0480,0030,0710,111

0,0130,0140,0240,0510,0870,1290,1810,214

0,0370,0350,0480,0730,1080,1480,1660,220

0,1020,1630,2360,3200,3960,4880,5190.526

gl*m2?

0;5860,9471,2861,5841,7361,700

———— 0,1250,2200,5510,8951,2151,5011,6321.567

— ci2090,317— 0,1630,3!30— 0,0910,4670,0120,5220,0300,540

— 0,1410,071-— 0,1380,149— 0,1210,237— 0,0930,299— 0,0440,3830,0100,4470,0750,4980,1170,501

0,0130,0410,0130,1190,0250,1870,0490,2680,0870,3420,1310,4170,1810,4600,2060,465

0,0400,0350,0360,1140,0500,1830,0710,2630,1080,3370,1490,4030,2000,4580,2220,468

).=0,16

–0,147‘— 0,1320,195— 0,1310,526— 0,1170,868— 0,0861.183— 0,047

0,0770,1450,2160,3090,3780,4470,4960,504

= 0,1280,1980,530

~ a= 0,20 0,8861,2071,4861,6201,568

— 0,1020,1880,4910,8031J021,3481,4781,429

1;4701,6201,561

—0,1140,1960.4890,8011,0921,3441,4801.410

0,0010,0630,107

0,0150,0120,0280,0510,0870,1260,1750.206

0;4360,4850,513

0,0510,1280,0960,2730,3390,4020,4540.486

-..— -0,0500,1190,1880,2700,343

– 0,1130,1960,4950,809

-.– 0,1180.195

0,0110,0080,0210,0500,0920,1280,1720,224

0,035-0,0300,0440,0700,1070,148

0,0540,1300,1930,2780,3550,4190,4670,487

0,0470,1230,1890,2700,3510,412

0;4830,8101,1001,3511,4961,441

-—.— 0,1060,1940,4900,7971,0791,314

a= 0,35 1,(W81,3571,4881,426

– 0,1120,1900,4820,7841,0581,3131,4551,419

0;4110,4510,464

0,0440,1140,1850,2680,3410,4120,4580,462

1,=0,55

— 0,1080,1920,4800,7871,0641,3061,4281,395

—.0,0370,0340,0480,0700,1050,1450,1990,222

0,0420,1200,1900,2680,3370,3920,4690,484

.-— 0,0980,1860,4800,7871,0761,3051,4501,422

1;4241,392

0;2060;4700,2510,464‘

cd

I?.A.C.A.TechnicalMemorandumno.

I

&-2

Figure I.-Notationof dimensiormfor table I.

to >Ca

“tQ8

Q6 i

0,4 A

- s-o- S-o,ie+ S-0,36-

0,2f

+a

2 4 6 8 lo f2 14 16

874

f,

a

al

a{

t7i

.0,2

Ub

Figs.1,6,7,13

Figure 7.-Polareof the wing withand withoutblower.

-.‘\

Fi&re 6.-Lift coefficientas p functionof the angle of att,::k.

,$,~&

N.A.C.A. Technical Memorandum No. 874 Figs.2,8,19,22

Figure 2.- Test.set-up for measur.ing the polars.

t3tream.Figure 19.- Test-up for the down-

wash measurements witha feeler vane.

ti.A.C.A. Technical Memorandum No. 874 Figs.3,4,5

\

\

(

Figure 3.- Cross section of fan and casing.a propeller, b guide vanes, c motor, d honeycomb, e mesh,

.

ltlt’ll ltltlil[~ l’”

ffi0>

0.l-l1

oN“

--n I!, !!!l!! l!.!t II

Iil+l+li 11

0.1+

0:

lHIHllllillH-0.l-l1

f-

m c%g

I0

0 u-l0E

X* . . . . .r-l d“ 1+

f throat./-

I

d

1..& cohrl-P.rno

z II

I

+“

- ---—--——.--.,,-—.7---..-.._

LAJ\-r+!Arq——

C3.

r-2,4f713417

hM57

,*

,+ P1,%0

.

P-1,500

\

+

x

?7

I

P1,916

. P2,417

1=-0/933

7.,

1

/

/’

Figure11.-Chordwieepressuredistributions.a = 12°.

.

IP

- s- 0,0.-,– s-0,18

— s-’O,36

~ .,*”

.FP-4,500 ~

f , ‘“m’

/-/,0 .

-2

-3,0

—~T

L?417

‘2,417

[-

‘1-1,916

7

/

m.’

lf67

rrr”rr ~~ “

1,s00 1S16 2,417 3,168 4,500

Y

. — S-QO

“%= -. –+– s-Q18

. s?036

(

~=,~(j

/

+

.

Figure 12.-Chordwise pressure distributions. d = 14.S0.

N.A.C.A.TechnicalMemorandumNo. 874

Figure14.- Incretnsein liftaccordingto the theory.Positionof

wing with respectto the propellerandthe relativewind direction.

s-0,f8 -—;

Figure15.-Lift distributionaccordingto theoryand measurement,s=O.18.

Figs.14,15,17,18

1 1

N.A.C.A. Technical Memorandum No. 874 Figs.16,20

.. .,

I \

!,:,1..,,

/ ● \●

●. . ● .f \ ●

9.

-- -- -—* — — -*— --- .1 ● 0,Je a, ~c,>,Q. .0-., ,..0 0.! .0 ~ -. .>> 12

/“/

-g -./

0

0,8 0 \

“ ‘-k● ● .

●●

s=== =m=u●

● -. -~ 9 8°Y- -.

e “

~-‘6 Ca

t “ ‘$

< ‘ ‘

.m q“! 0

t ‘

*F- -=-- -

--$0,2

---— Compufcdby Lofz[ WL9

0 Measured I 010.7,

Compufedby .toniriq~

● ‘ Meu.5uredJ “nq~z~~hn

‘6 ‘5 ‘4 ‘3 ‘2 ‘f w 123456

“Figure16.-Lift distri-bution accord-ing to theoryand measure-ment s= O.36.

S-0,36—i

VRnehd

4’- Jt0,25 t

1,5 t 0,5t

N.A.C.A. Technical Memorandum No. 8’74 Fig.21

,, -,

~s= o+“ s = 0.18x s = 0.36

60 ~.

, /f-

8 - //+

6 /.—

/.x/ h\ \

4 u/‘ b//

2 y“a/t = o /’

aft = 0.3 21/

8

/36 ~

//’

*9

//~ \. ,“

4 - \

/.. {

,*/2 , /0’

aft = 0.715 * a/t = 1.1 25*-

/

I0 3 6 9 12 15 0 36 9 12 15 a

Figure 21.- Variation of the downwash angles.

I I mmmmmmn- ,,...-,...-- .,,, , .,, ,, , ,,.. .,,., ,,,.. .,,., , , . . ..... . -,.---.,-. ..——

N.A.C.A. Technical Memorandum No. 874 Figs.23,25,26

\

yi

~7 I o-

1 II I M S

, II c

+~___ --—— —. —— — -t-

.14_— .

I ggI

In.#

-J_t

1

c<a

3 6 9 12Yl% Figure 26.- Wing lift in

Figure 25.- Spanwise distribution of propeller slipstream.dynamic pressure in propeller slipstream.

..--!.............. ,.,,,,,,.,,...... ,.,. ,..,,____

15 a

M.A,C.A.TechnicalMemorandumMo. 874

iv —

l%

IO.a-

0/5—

0,+—

az—d~

ij

JL

az—

ai I a2

—

Lj“—

—

I

—

——

—

&

Rc-“

afz a76

- c-a3 0,?

/-Figure27.- Polar with and without..’ propellerslipstream.

,

Figuro 24.- Spe.nwisevariationof jetrotationangle in propeller

slipstream.I<.?I

i

do

tr \f6

1?

, i 9

1

- f,o 0,2 0,4 L16 0,8

—k

-4

.8

— +=Q25

V2— X“=0,5— fl=f;o—-0-- ll=l5— 11.20A 44 = 2,75

16

-P.=”!+.

>. -“’==!!!!

Figure 28.- Chordwise pressuredistribution,a= 4° zo

Figure 29.- Chordwisepressuredistribution,a = 8? %

w ~Tw %0 P,

‘PPf,[0

-3,6(XI : -25.?.3~ i

o

/.

.1,0

-2,0

- ~o

+“

~f

\\ Q333: 0.6s7

r-!933 r-1533 r-[ml P-0933

b-0,567

k- Q333

/“

. ..

k0,000\,

,/,/ :

(,.

FF~~2,533 WL13

.-Withoutprop.—With prop.

.,d=120:

1 ,,

Figure 30.- Chordwisepressuredistribution,a= 12°

---

r..r rr.r d= 14,90

w

Pk

2..

Figure 31,- Chordwisepressuredistribution,a = 14.9°

\ II I I b L’w I h

1 1 1 I 1

/4 l~8-t---

IL 1 1 1 I

I “Ft

FFwlI I 1 1 I

I}q.4”

II— Wt”fhou~

— Wifb $+j’ —;A

-5 -4 -3 -2 -f /2345

{5

O,f

Figs .32,33,35,36

4“ 8“ 12°14.90

— + =@-33 -:- -Q533

rlgure 36.- Downwash bebindFigure 32.- Spanwise lift distribution. the wing in absence of mco-

fiellerslipstreati.Vane

d

PropellerWing ,

a.—._ ._._ —.-—.—.—._.______

Motor

tQJ6 t

f.t 05 t

Figure 33.- Arrangement of apparatus for downwash measurements.

Az#?effsurtng

p/Qne \z

/\ ~:- ----- ---_—-1

1

,,.. 3 1—.

:.—.— .— .— .— .—.— ——.— .—.—. -—. -— —- .—1-.—.— —-—

x IL -J 1

i 1

L--—--- .--- __.. J

1’ 1,5t f0.375 t T-z

Figure 35.- Position of measuring plane.

N.A.C.A. Technical Memorandum No. 8’74 Fig.34

63

8

6

4

2

0(

8’

6

4

2

(o

~Without propellerx With II

6°

8

6

4

2

0

6

4

2

0

0 36 9 12 15 (] 3 6 9 12 15 a.

I?i=gure 34.- Variation of downwash angle.

....6

. .

N.A.C.A.TeohnicalMemorandumNo. 874 Fige.37,38

z—.

-3.0 -2,5 -2.0 -1,5 - ID -0,5 0,5 f,o 15 2,0 2.5 3.4

Figure 37. a = OO.— d“.y

I

C(=o”‘R

f.,

I I 1 I I I I-? I

2,133

III 1“ I1: !,600

I I

1,067

0.633~. –

t

0,593

\ I /

I I I ! is 4

! !-1,087

{ J I !

I I -1-- L I [ I 1 I

! !- Z,fss

I 1

-ao ‘2,5 -2,0 -46 -1.0 –0.5 0.5 f.o i6 2.0 2,5 3.0.“

Figure 38. d = 4°.--d”—-y C(=4” —%

Downwashangle and dynamicpressuredistribution.

11” —

N.A.C.A.TechnicalMemorandumNo. 874 FigB.39,40

,.

–30 -2,6 -2,0 -1.5 - !,0 -0,5 0,5 f,o f,5 2,0 2.6 2,0

Figure 39. ci= 8°. ~I&,

e

Figure 40. u = 12°. ~ ‘Mq.

Downwashangle and dynamicpressure

—id = Iz”

dletribution.

.

N.A.C.A. Technical Memorandum No. 874 Figs .41,42

+

2133

tao

[067

0533

:“

@33

1,067

1$00

Zfm

-30 ‘2,5 -2,0 - f,5 -1,: -0,5 0,5 (o 1,5 2,0

—. d“—$ 30

—-y d = 14,9°

Figure 41.- Downwash angle and dynamic pressure distribution.

I

I

.

MotorI

1

Figure

.,

42. - Xodel wing with propeller.

N.A.C.A. Technical Memorandum No. 874 l?igs.43,44,47,7?

Figure 43.- Wing with motor not Figure 44.- Wing with motor enclosedenclosed in fairing. in fairing.

Figure 47.- Bearing pieces forpropeller shaft.

Figure 77.- Test set-up for the down:wash measurements with

dynamic pressure sphere.

—.. ... .. . ... ..——

—

N.A.C.A. Technical Memorandum No. 874

. . .

Ca

1.6

n-ck

/’(1.2

/c(

.8 c

. .

c

.4 /

/

o @

-8 0 8 16 aFigure 45.- Lift curve of

wing alone.

rniigs.45,46,48

Ca1.

1.

.

.

0.2 cm 0.4 0.646.- Polar of wing alone.

,//“—\

/’

Figure 48.- Position of propeller aXiS K= 9°,dotted curve K= -6°.

N.A.C.A. TechnicalMemorandumNo. 874

r,

Fiare 50.-Characteristicsof freelyrotatingpropeller.

~~Oz ,16J

., Figure 51. R = 9°.

Figure 52. K = 4°.

Figure 53. K S=-lO.

Figure 51,52,53.-Effect of propeller slipstream on lift coefficient.

I111:;

N.A.C..A.Technical Memorandum No. 874 Figs.54,55,62,63,64

rr’r

!

t-t//~+

—a

! I,~.q-

Figure 54. K =–6°.

-—?0

y ‘8u

{4

uy

to h /“

A.413(‘q=,to)

-a-#” -5 s m 13 i%’

— X-9”

J w — X.4;— x.-l— ~..6W

‘Q4

Figure 55. A =0.13.

Figures 54,55,–Effect of propellerslipstream on lift coefficient.

Figure 62.-Polars of wing in presenceof propeller.

Figure 63.-Thrust coefficient asfunction of A.

-J

C.) .?.0~a

t28 r

).(6 J ,

14

a

20x-90

‘8

Q6 h

3 I%

A-L@ Z= :W

Q4— A-QU”~= ,00— A.ox T=-=,26— A-Q* T=.,JU— A-Q13T

02—----W/n=&” ‘Eo

— cm 4-1%

J__LLT’

Figure 64.–Moment curves of wing in ~presence of propeller.

.

N.A.C.A. Technical memorandum No. 874

0.13

0.16

0.20

0.350.55

iCaFigure 56.- Lift slope ~ as a function of K.

Figs.56,57./

5.8,’~idca,.,.zr-.

..‘\ 5.4tk

Ff,?] c—————K =40

5.0 — +— K=-low’

,.

4.6

14.2.

‘+

0 ● ~cl .20 .30’ ,40 .50’ ●GO A .<:.,.,;m,:t‘.,-rLdca

Fi~re 57.- Lift slope — as a function of A.do

N.A.C.A. Technical Memorandw No. “874 Figs.58,59

Ca

2.0

1.5

1.0

.5

16”

.—

I I I I I J

-.5 -.4 -.3 -.2 -.1 0 .1 .2 .3 Cw

Fieqre 58.- Polars of wing in presence of propeller,K= 9°.

1.5 -

1.0 /4’ —- ___

7

/1— —Wing alone

.5 ‘~ ‘–– – –~ -—+”f-–

J/’_ X—A= .35

+[+—A= .20

--–-+- – –--x u—-A= .16

0 +-~ = .13[ \

-8° --––- - -L- - ‘-+- ‘– xh

-.5 -.4 -.3 21---- 0 .1 .2 .3 Cw

Figure 59.- polars of wing in pre$ence of propeller.K= 4°.

II 1111111ln-lEn- Immmmmmllm ImIIImII I .,.,, .,..-,,, ,,, 8 8 , , , ,..-. . . . . . . . , ,. ,- , .,

N.A.C.A. Tectiical Memorand~ No. 874 l?igs.60,61

.

Ca

—.-

1.0

k

2.0. 16“ I I 11, 1

1:2°

1.5

,-!/c

0

-. 5

A----

-n &-~+--+---/ -. -..-k

-— _ ~_-+-. _.

--,?.

--— /./ I——Wing a“ one-1-yj’

0“+ .-–. ._&. .__~ -_/ -f–. -–+_ ___ {j ~~ =.161

1 1 ‘[MA =.13

-–—– –~–– – -+”–‘—–-X

‘ L-L-.1 .2 .3 Cw

L-.-l__lL_lL_l I.5 -.4

——-.3 -.2 -.1 0

~-x =.55k— i =.35+— A =.20

Figure 60.- Polars of t’iingin presence of propellsr,K= -1O.

Ca

2.0

1.5

1.0

.5

0

- ——!

—

—

~A= .16~A= .13

I-8°~ -–- –A–’ -–+–l---- ‘-L

.—-. 5 -.4 -.3 -.2 -.1 0 .1 .2 .3 Cw

Figure 61.- Polars of wing in presence of propeller,K= -6°.

1,-:11’!’

N.A.C.A.TechnicalMemorandum Ho. 874

-’6F+=H+7’

I

++zf!$f!za.I 1

L5’z/l/;02 03 a4 ‘“”Q!I I I

O.A “ “(i I J

1.6

1.4

1.2

1.0

.8

.6

.4

.2

0

-.2

?Igu.re 65. K =4° ~

c. P I I I I P I-= LLt

/- I I

I

I Am ~/

,xV//l I I- cm J#~,/ fl I -; “1 “~

Figure 67.K=-6° d

Figs.65,66067,68

L

L(

.!

o

:5

Figure 66. K=–l” w’

Moment curves of wing in presence of propeller.

.,. ,.-—,

N.A.C.A. Technical Memorandum No. 874 Figs.69,70

“ilcm ‘“ -.

dca

,,

. ,.

&4+

- GfcL

.27]I

I I I 1’

Wing alone ~=~)—-— -—

.25 l--

/ <1

/ 0/“

.23— -—-

)(

/’f- -~

I

.21 “

.lYL ‘kl

I

‘--- , .Z10 .2 .4 .6 ,8

dcmFigure 69.- “ “ . .

::a::;;:;l::efflclent ~of wing in presence

●

dcmT ~ “ ‘-—-

-—. .

1.2-— I

?I ,—o—~. cJo

.8 x ~= 40+— K= -10A-K= .6°

.4 -

dcmfiiguxe70.- Variation of ~ with ~ an:i K.

i’z’u

ta

!6

14

./-.”......... lo

*

Q8

Q6

h

H&....--.-””””

+--02

$~h

-U -Jo -<5 -2,0-TJ -to -qs 0s !0 t~ Zo 2,5 3.0 45

FRw$J!i ● ‘.’.1

..-

m m I 1“’”1/-l\l 1“ I-A”OX I I

I7A .1 F&l I Ia.-3”lI I*

........’”k

>jt~ 1111111:11X.40

–+ “k~$

lLLL/lYTN ! !---::”ntntI“’- ... .,,,,.,

=+

I I 1 1 1 1 1 1 n ! 1 I , I

-35 -30 -2) -Zp -f.3 -to -qs Q5 to 73 20 2s 3.0 s

.

Figure 71. K = 9°. Figure 72. K - b“.~

$

Spanwiee lift distribution..

z.

2

—

N.A.C.A.TechnicalMemorandumNo. 874 Flga.73.74

(8%

I(6

i

...............- ffl

+

Q8

Oi

—04

.

.-.-’ ”””z -

?J$

t~

-$5 -30 .2s .20 -f.5 -lo .05

Figure ?3.

11~ 1-----------------l---- —!tttmtT’FtH

%

o

-: ___~’OrTT-iT‘:--”””I I

........-.””””

$G>

“P -1? ‘2S -Zo -v -20 -Lv Q5 10 !5 213 2> 3B 3J

Figure 74. K = —6°.

Spanwise lift distribution.

. .:,<.

.-

,,...-,......J ..

m

—

. . -“’”””1

%$!6$“U -jo -25 -2(I -~ -!0 1

I Figu-e 75.-Spanwlse lift distribution.h= 0.16I

(1

km%

*K%

u ‘b !0 I “—u -a

WH *N

Figure 76.–Spread of flow Reparation overthe wing.

.

y.,/

N.A.C.A. ‘TechnicalMemorandum No. 874

_—- - -- ,,-.—.-----”

- —------------ ‘“-

*TWing$4 =S?:e=em Position

- o“ i -----

A.....

— t...... ..—--””---=-Pos~tion

2.5 t-—q

II

I

. I

Figure ‘78.-Arrangement for downwash measurements.

6

12.0°

8.0°

4.00

0

12,00

8.0°,.

4:00

0

-2.

Position I

Figs.78,79

a

16°12°~o

00-4°_sO

160

12°

4°

0°1 1 l\ 1 f 1 1 #

-1J o

i ‘“4

-+80gt’ --0 -1.5 -1.0 -.5 0 .5 1.0 1.5 2.0 -k-

I’igure 79.- Downwash behind wing without propeller.

—.- -— .—

Fig.80

Figure80.-,Dowmash distributionwith pro~lleriK=9°, u- .SO

IT.A.C.A.TechnicalMemorandumIio.874 Pig.81

k-Q3

1 I I

,.

Figure 81.- Dowmash distributionwith propeller.IC=90,Cf=8°

o

,PositionI , ‘-W

o

A- 035I I i I I1 -i

---

—— —

,i

1I

I?1, I I I I

,I1 i-am‘0

I I I an

2° I I I I I wo & *o I 1: I-3,0 -2,0 x.-69 -to

-2”’LX--3* 20 warn

‘# ‘~—- ~

K=-6°,se-3°

Fimre 82.- Dounwashdistributionwith rmoneller.—-—. —.-— -—- .— ● ✎

,*0

‘k9

M.A.C.A. Tocklnic81

I

Memorandum Ho. 8?4 Fig.83

—

I I I

R

Figure 83.. tinnw418h di8tributi0n with propeller.Ic=-60,U=8°*

I

li.A.c.& TechnicalMemorandumEo. 874 Fige.84,a5

L.,.!1

-;

u ,.

,- ,. ,

!5 -w -0

—~

2P+ A-QA!~ A*QW

- A-q.m:-7-L.Anqafolu

PositAn

x

Figure84. K=9°, a=8°

~ i4

b-.X. -(j o

-25- a- u -0 Q.s Y7 @ 10 .!5 2p

+ x-Qm— A-@o

_$ ?

+ .i-03.5+ A-Q.%----wiq llfm-le%s.0.

/ ~ —.- —---.-------- ------ ..-,.

,.

Figure05. K==’-6°+.a=cB”

Spanwise variation’of~, the ‘tq\l flow e$ficiencyn.

— _..

. ....

~..*’.

i-F

f