Experimental investigation of the aerodynamics of a wing in a slipstream

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OF A WING IN A SLIPSTREAM

APRIL, 1957

BY

M. BRENCKMANN

Bibliotheek TU Delft

FacuiteH der Luchtvaan· en Ruimtevaarttechniek KJuyverweg 1

2629 HS Delft

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--,,--.-"""""'"---:---

- - -- -

-,

b,;-,liotheek TU Delft

Faculteit der Luchtvaart-en Ruimtevaarttechniek Kluyverweg 1

2629 HS Delft

EXPERIMENTAL INVESTIGATION OF THE AERODYNAMICS OF A WING IN A SLIPSTREAM BY M. BRENCKMANN

fl'fil'jiliiï

JII

De 1 f t /LR C 1955419

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..

The author wishes to express his thanks to

Dr. G. ~. Patterson for the opportunity to pursue this investigation and

for his interest in its progress .

The direction of Dr. G. K. Korbacher and Dr. H. S. Ribner

is gratefully acknowledged.

This work was made possible through the financial and

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'

.

..

•

SUMMARY

An experimental study of a wing in a propeller slipstream Iwas made to determine the distribution of the lift increase due to

slipstream at different angles of attack of the wing and at different free stream-slipstream velocity ratios, the results being intended as an evaluation basis for different theoretical treatments of this problem.

During this study it was found that the main departure' from the potential flowexpectation occurs at large angles of attack, where a de stalling effect of the slipstream on the whole wing results in

significant additionallift. Measurements of center of pressure shift were made to clarify this point.

In other respects potential flow representation of the phenomenon seems justified, provided slipstream rotation is taken into account when solving for the spanwise lift distribution . In terms of total lift gain, a closed form solution given by a slender body theory showed good agreement with experimental values from which desta:lling effects were s ubtracted.

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TABLE OF CONTENTS Page NOTATlON ii l. lNTRODUCTlON 1 Il. APPARATUS 1 2.1 Wind Tunnel ₁ 2.2 Wing 1 2.3 Motor 3 2.4 Propeller .j3 2.5 Pressure próbes 4 lIl. PROCEDURE 4 IV. CALIBRATION 5 4.1 Wing 5 4.2 Pressure Probes 7 V. REDUCTlON OF RESULTS 7 VI. ACCURACY 8

VII. RESULTS AND DISCUSSION 9

VUl. CONCLUDlNG REMARKS 12

REFERENCES 14

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c d Pl P2 PT

J2s

Ps

q qo .9.j

tt

_J Kq A qO·+ qj q

=

2 y

l

A

Crw

( ii ). NGTATION wing chord

distance between straingauges on flexures pressure measured at yawmeter manometer pressure measured at static probe manometer total pressure in the flow

static pressure in the flow

static pressure at test section wal! dynamic press ure

dynamic pressure in the free stream dynamic pressure in the slipstream

equivalent dynamic pressure (average of qj over

slipstream cross section)

equivalent dynamic pressure increment (average of qj - qo over slipstream cross section)

"average" dynamic pressure inside and outside slipstream

strain measured by straingauge 1 or 2

strain measured by straingauge 1 or 2 during calibration

spanwise coordinate. origin at midspan 1

distance of the center of pressure aft ofTchord point

lift per unit span

aspect ratio of immersed area

gauge constants • bending moment/strain local lift coefficient of wing at midspan in

unif orm flow )

local lift coefficient. non -dimensionalized with

A. the average q

(7)

D L AL p R

s

T

v

.

J

cx:.g

Ó

=

tan -1 DIL

cr'

=

ëij -

qo qj - qo

locallift coefficient. non-dimensionalized with the freestream qo

drag lift

lift increase due to slipstream

"potential flow" lift increase due to slipstream aerodynarnic moment acting on flexure

aerodynamic force acting on flexure propeller radius

area of the wing immersed in slipstream propeller thrust

free stream velocity

average slipstream velocity (based on-ql

.r

angle between chordline and flow

angle of attack

angle of rotation in slipstream (helix angle) angle between flexure and flow

glide angle

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( 1 ) I. IN TRODUC TION

-Interest in-8TGL -a'Ï:r-craft -configurations- has put new

-·emphasis t'm"propeller--s1ips-tream deflection-as' a means-·-te create lift at low-t'-ofWant·velocities. Theear-lier theories-

-<Rcl,.

1 and 5) are knowrr to-'be inadequate -ror 'practieàl-eonfigu-raHorrs -arrd-the validity of

"recent theori.-es -,(Refs. 2-, 3-, 4) is -not defi:nitely~stahlished-.

-Experi-ments"on-this subject'fnot aUlisted) are-not·sy-stemat±C from our point of-view -m.ainly because they were 'conducted for ether purposes .

In fact only-R:eferences 5 -and 6 a're -cited for'cmnparison with theory

{Refs. 1 - 3). -Furthe-rmore the circumstances of their experiments are not too close to those'-encountered in .. STOL aircraft (weak jet,

small aspect ratio of the immersed part of the wing).

This investigation is intended:

1) to supply experimental data in the form of the spanwise

lift distribution of a straight wing in a propeller slipstream

for various angles of attack and for various freestream-slipstream velocity ratios,

2} to make an attempt to separate the different causal

relationships combined in the creation of additional lift,

3) to isolate important parameters which should be

included in theoretical analysi~ .

In this experiment flap action of STOL applications will be simulated by wing incidence to avoid mechanical complication of the

model. The results wi11, therefore, describe absolute lift increase at moderate flap deflections only, but its rate of change with angle of attack could be interpreted more generally i. e .• for large flap deflections as we 11.

II. APP ARA TUS (See Figs. 1 to 6)

2.1 Wind Tunnel

The UTIA 32" x 48" low speed wind tunnel was used in this experiment. The lower speed range allowed a freestream speed

variatic)fi from zero to approximately 100 ft. /sec. The wind tunnel balance was not used.

2.2 Wing

The determination of the spanwise lift distributinn was the main object of this study. A wing was therefore needed to allow

detailed and rapid sectional lift measurements. The direct measure-ment of lift force by multiplestraingauged elemeasure-ments was chosen and preferred to indirect methods as e. g. the measurement of pressure distribution on or in the wake of the wing.

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Wing-tip and wing-planform influences were assumed to be of secondary interest. Therefore, a wing of constant chord and

co~tant profile was designed to span the tunnel vertically from wall to wall (span 32 "). In order to avoid excessive tunnel wal! influence the chord to tunnel "wid.th" ratio was chosen as 1/6 (chord 8"). The profile used was a thick, laminar profile (NACA-64 A 4 18) as used for wings of slow transport aircrafts.

The wing was composed of ~see Figs. 2 and 3}:

a) a spanwise spar held by pivots at both ends in the tunnel walls. lts pivot axis corresponded to the 1/4 chord axis of

the wing a.nd intersected the tunnel axis at right angles. The pivots allowed for variation in angle of attack,

b) 19 flat aluminum flexures .052" thick, attached at 19 spanwise stations to the spar, extending towards the

trail-ing edge and forming a 7 1/20 angle with the chord line.

They were free to deflect perpendicularly to the wing plane. Two SR4 AB straingauges per flexure we re bonded. 5/8" and

-5/8" aft of the 1/4 chord line so as to measure bending strain,

d

19 laminated wood wing sections attached at the free end of each flexure. These sections were hollowed so as to enclose the spar and the flexures all but for an inspection

opening covered by an aluminum plate on the underside of

the wing. Each section was separated from the next section by a l/S" gap and thus was independently

trans-mitting its share of the lift through the flexure to the tunne 1

walls.

d) a plastic skin wrapped around the whole wing. taped to

the trailing edge. sealing the gaps between the sections but

transmitting very little force from section to section . This skin was made of .002" SPRAYLAT , reinforeed with

adhesive tape.

e) a wiring syetem enabling each straingauge to be placed

individually in the bridge circuit of a Baldwin SR4 strain

indicator (see Fig. 5).

A temperature compensating dummy gauge was placed on the spar in a non-stressed position. The selection of one particular straingauge for reading was achieved through a multiple rotary switch (for the 19 span-stations) and through a double throw switch ([or the two gauges per station). The difference in strain between these two gauges was the measure for a load applied perpendicularly to the "f1exure ~

lndivudual straingauge reading was necessary because of variation in effective gauge factor from gauge to gauge.

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{ -s )

The 19 sections of the wing were, divided into 5 groups: 3 elements (2" wide each) facing the core of the slipstream, 3 elements (1" wide each) on each side facing the slipstream boundary, and 5

eleIp.ents (2" wide each) on each side facing the freestream (see Fig.

U.

The top Pivot of the spar was arranged to have stops,

setting the angle between chord line and tunnel axis at 00 _, ₅0

, 100 and

150.

2.3 Motor

In <?rder to drive a propeller of about 8" diameter, shaft speeds of 10,000 to 15,000 rpm were required. Electric model motors of the variabie frequency type were not available and trans miss ion problems made the use of convent~onal electric motors difficult. The adaptation of a pneumatic hand grinder motor was suggested and proved

to be a good solution . '

In this experiment the v,ariation of the freestream slip-stream velocity ratio was achieved by varying the tunnel speed and keeping the air pressure for the pneumatic 'motor çonstant. Fig. 7

illustrates the power characteristic of the motor for various air pressures and the region of operation (0.5 H. P ., 13,000 rpm).

The aerodynamic shape of the motor wa$ found satisfactory "as is", (2" max. diameter, 12" length), so,that.it could be used

without nacelle, and put inside the tunnel, ahead of the wing in a "pusher" configuration. It was fitted with a 1/2" air pipe 4' long extending into the settling chamber of the tunnel and suspended on 8 piano wires spanned diagonally to the 4 fillets of the tunnel throat. The axis of the motor coincided with the tunnel axis (see Fig. land 4).

The compressed air supply was delivered by a mobile compressor unit of 70 ft. 3/ min . capacity . A rubber hose connected

the compressor to the air pipe of the motor through the bottom of the settling chamber. The motor exhaust air 'was cb.annelled into a conical

fairing along the shaft extension and expelled through an annular slot around the propeller hub.

The supply pressure of the compressor was set to be 103

to 105 psL and stabilized by a discharge leak. 2.4 Propeller

Since themotor was arranged t<? drive a pusher propeller and had a right hand shaft rotation, t~e propeller had to be left hand

rotating. A series of right hand rotating model propellers we re tried out in bench tests with the pneumatic motor. A propeller of 8"

diameter and 8" pitch was selected for giving the highest average velocity of about 90 ft. /sec. It was copie' into the left hand rotating propeller made of laminated wood and fitted,to the motor shaft with a

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2" hub cone. During the experiment the propeller disk was set 6"

upstream of the 1/4 chord axis of the wing.

The static power characteristic of this propeller is shown

in F1.g. 7. Later measurements provided the thrust characteristic shown

in Fig. 9.

2.5 Pressure Probes

In order to obtain total and static pressure measurerrents in the rotational slipstream, a special combined yawmeter, pitot and

static probe was built (see Fig. 6).

A .75" diameter steel tube 48" long was introduced

horizontally into the slipstream through a wooden fixt ure replacing the

tunnel window. lts axis was arranged to intersect the slipstream axis

at right angles. Radial traversing of the slipstream was obtained by

sliding the tube in or out manually, measuring the travel against stops.

A .625" diameter alu.minum cylinder 10" long was mounted coaxially at

the end of the steel tube. Along a perimeter of this cylinder, 2" from

Hs tree end, th."fee . 020" diameter holes were drilled radially, with a

450 _{angle between the}_{central h}_o_le_axis_a_nd_{each of the two other}_hole

axis. These holes were connected through plastic tubing running inside

the steel tube, to manometers. The central hole was used as total head probe and was connected to the pressure side of an ethylene glycol

manometer. The two otheJt' holes were connected to a zeroing U

manometer. The free end. of the aluminunl cylinder was equipped with a small statie probe made of .048" diameter steel tubing bent into the

stream direction 2" away from the cylinder end and aligned with the

middle hole. The portion of the tubing parallel to the flow was . 875"

long, and had four .030" diameter holes .250" downstream of the tip. The tip was made of a hemispherical brass plug. The static probe was connected through plastic bbing to a second ethylene glycol manometer. The outside end of the steel tube had a comparator disk with a pendulum

attached indicating the angle between the total head (or the statie) probe

axis and the vertical.

Zeroing of the U manometer permitted aligning the total head

probe hole axis with the flow. and measuring the total pressure at this

point. Reading of the comparator measured the rotational angle.

Keep-ing th is angle constant and placing the static probe where the central

hole of the yawmeter had been, yielded the static pressure at this point.

Tunnel flow speed was controlled by measurement of the

pressure d.ifference between test section and settling chamber with a

water micromanometer . lIl. PROCEDURE

Preliminary straingauge readings were taken to assess the sensitivity limit of ,the wing, which was found for Q{.~ = 00 about at

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qo = .2 in H₂0. At higher tunnel velocity and at

<X-" ,

= 150 , the wing

showed a tendency to buffet, which limited uSefu1.measurements to qo

L 1. 7 in H20.

Investigation of the slipstream in the streamwise direction showed that rotation and velocity changed very little withinthe first 10

'1

downstream of the propeller disk. The distan,ce from propeller disk to wing 1/4 chord line was chosen arbitrarily at 6".

The test runs were made under the following conditions :

Spanwise load measurement no slipstream

(W ing alone in tunne l)

Spanwise load meas ure ment with slipstream

(Wing and propeller in tunnel)

Slipstream cross section (Propeller alone in tunnel)

( ( ( ( ( ( ( ( ( ( ( (

Angle between chordlineand flow:

Tmmel-dynamic pres8ure: ,

"

,.

.

. .

<IQ

'

= 1.70'-, 0.

75t:

,

O.42 ~

, 0.20

~

in H 20

Ahg}e'be-tw~r ,chordline and flow:

. . • IJ·

0( .,;. 5° 100 150

g.

'

,

( . Tunnel dynamic pressure:

( ( ( ( ( ( (

qti

,

.;

'1. 7 èÓ{.~:· 0 . 7 5 c, 0, 42' , .0. 20 '

an~ ~ ~ 08\'

~

_{(tunnel-bff) in H 20}

Tunri.el dynamic pressure :

In the tunnel~off corîdition the propeller sustained a weak reci~culating flow in the tunnel, so that exaGtly".static conditions were

not obtained. '." ' .

Every spanwise load measurement was followed immediately by a zero reading of the 38 straingauges with:tunnel and propeller off. Calibration of 'the wing was performed twic'e wlthout the plastic skin before and af ter the experiment and twice w,iththe plastic -skin during the experiment. The yawmeter calibration wassinlU~taneous with the

slipstream measurements. ' .

IV. CALIBRA TION

4.1 ,Wing

Each flexu,re behaves like a cantilev~r beam with a normal load Pand a moment M applied at its free end~ , 'p, andM are here the aerodynamic force and moment transmitted by tJ'le Wing section to the

" , ' .

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flexure . The distance of the straingauges to the point of application of

the force being dl and d2. the strains sI and 8-Z-measured by the

straingauges can be written

Cl 2 = Gauge constants

• (bending moment)

( strain )

dl - d

2 = d = Distance between gauges.

which gives us the reduction equation:

If during calibration we arrange to have s2 = O. we obtain one of the constants in the reduction equation as: _1_ = l a n d similarly for

CId sc1

sI = 0 we obtain the 0 her constant as: 1 -P

C 2d =

~c2

,

-The wing was set in a fixture so that the flexures were

horizontal and the wing sections free to deflect. One section at a time

was loaded with known weights placed so that one of its two gauges

indicated no strain and other gauge indicated Sc l' This procedure was repeated for the other gauge. giving sc2'

The calibration was repeated for different weights and the resulting s 1 and s 2 plotted against P. The slopes of the curves thus

obtained

w

~re

takerI as the constants in the reduction equation .

For the later discussion of the results it is of interest to

know the point of application of the load on the wing sedion, i. e. , the

center of pressure . lts position, expressed by its distance downstream

of the 1/4 chord point, follows from:

s2 SI

- -

-x =.625 C2 d - 1. 625 CId (in)

cP s2 sI

C 2d - CId

Since small deflections actually occur under load, the

stretching of the skin over the gaps between the sections was expected to

have an influence on the readings. rCaIibr&tion with the skin on the wing showed transmission of load between the sections which reduced the slope of the calibration curve by an average of 5%. This reduction

however, would apply as 5% of the load difference existing between adjacent sections . This amount had negligible influence on the load

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( 7 )

Aging of the straingauges detected between the start and the

end of the experiment amounted to about 5% and was taken into account

when' reducing t~e data.

4.2 Pressure Probes

~ comparison of the tunnel's dynalnic and static pressures

as measured by the tunnel probes and by the yawmeter for different

speeds led;ttêf''the reduction equatioIlS: '

P - P

T

s

= P 1 - .01 .995 P s - P s

=

P 2 +, . 0 1 q q q = P -P

T

s

= _.P ₁-.01 - -P2 ' - (in Ethylene Glycol) 1.01

v.

REDUCTlqN OF RESULTS

,

With the reuuc,tion 'equaiiun-the-straingauge -readings were

converted in1lo fbrces normal to the flexures. These forces were taken

as an approxima-tion 6f

tne

~

tt

within

our range"ot-'ac-curacy. This is justified. since I \l ,~',:,:.'

cos(

cf

-

r'

)

P

=

L cosJ

0=

P

=

angle between flexure and flow.

varying between t 7 1/20

local glide ,angle of the wing.

tan- 1 DIL. varying fr,9m 5 to 150

Results were plotted ás lift distribution curves

l

vs y

for the different conditions with and without slipstream, The total lift increase d'le

,.

tÜ

the slipstream wasobtained by planimetry as:

AL

=

,

f

(

l

with

span slipstream

l

withou t ) dy

; slipstream

R.eductionof the pressure probe readings showed;that the

dyname pressure of the slipstreJUn, Yias far from a constant incre'ment

compared with 'freestream'. An equiválent constant increment was

therefore defined as: .'

2

=

2 R

J

R Y (q. _ 'IL) dy J ,. , , 0 '

(15)

form, a reference quantity which does not disappea:r for either of the extreme cases - no slipstream or no freestream - was chosen as the

Haverage" dynamic pressure q and defined as:

~ qo

+

qj

-q = 2 with qj

=

qo

+

Aq

which means that when qj ... qo (no slipstream),

q

~qo

and when qo ~ 0 (no freestream) ,

'q ..

qj / 2

The non~dimensional quantities used are then:

local lift coefficient :

ê

L

=

l

/~

,

{

non-di-mensional stipstream strength

<ij -

qo

() =

qj

+

%

=

~~

This particular parameter 0" arises naturally in the

theoretical formulation in Reference 4. The results of slender body

theory {Ref. 2) are also reexpressed most simply as a function of

cr' .

The more common coefficients CL = '/,/qê-are easily calculated from

those used herein as: ê'L ~:r

C L =l-G'

Total lift increase values were rendered non -dimensional in

the following formS as explained in the discussion of the results ~

AL

ALp

ÀLp

ALp

q' CLw,S

~q' ex.

T' 2'rrOlS

~q'

C

Lw

S Aq

R

where the thrust T was obtained as: T· 41T

J

(qj

-'h

_j

q~

)

y dy and

o

C ~XT is taken in the center of the wing, in the normal flow without

slipstream.

VI. ACCURACY

The accuracy of the force measurements was determined by

the readability limits of the straingauge indicator: t 2 micro in/in.

Four strain readings we re necessary per sectionalload. so that an

average point scatter of

±

.02 lb/in was expected. There were-enough

points (18) along the span in order to detect this scatter and obtain a

reasonably faired curve. The accuracy of the curve itself is then close

to t .005 lb/in.

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.'.":' '~'. ' .'

( 9 )

. ,

estimated to be t .01 in EG L, which determines the accuracy of the

following quantities as:

2

q:

t

.

00 1

Ub

I

in )

T:

±

.09 (lb)

Converted to their final forms, the accuracy of the different

quantities is:

( point s-eaHe-r:

C L (

r.03 for <3'

=

.2S i"'--.1 for

G-

.::

=

.

\'88

.. ( faired curve:

t

.01 for G'

=

.26 +.03 for ( J

=

.88

AL

lift increase: t .2 (lb)

AL non ~ime'nsi'Onal: ·t 10% for cf

=

.26 ·t 30% for

CS'

=

.88 VII. RESULTS AND DISCUSSION

Figure 8 shows the dynamic pressure profiles and the angles

of ratation of the velocity about the tunnel axi~. along a diameter of the

slipstream, for the different conditions investigated. The overall

rotation in the slipstream decreases with inereasing f;reestream

veloeity Hncreasing advance rati<;l). The diameter of the propeller

eoincides witb a snarp increase in dynamic pressure in all cases;

there is no contraction noticable . The wake of the motor in the center

of the slipstream waspartially "fil1ed up" bythe m.otor exhaust air

expelled around the hub. The diátribution of angle of rotation .could be

closely approximated by a constant angle along the radius.

Figure 9 shows the thrust and equivalent dynamic pressure

inerement computedfrom the previous results as a function of the

slipstream strength parameter (S. Thrust and ~ characteristics

are s lightly different. reachinga maximum at different

0".

This wil!

be reflected in the nondimensional expressions for the total lift

increase. On .this same figure the meaning of

a

is illustrated by the

curve V o/V j vs

cY

.

A --, -:: Figures 10 a - e show the spanwise distribution of lift as

CL vs~/R curves. We can note the following

points:-"

a) the rotation of the slipstream effectively reduces the

lift gain in its downwash region and augments it in its upwash

region. The maximum lift isfound shifted from the slipstream

axis towards the upwash region . This Jnfluenee diminishes

with higher angles of attack and seemS compensative in overall result,

b) the limits of the direct slipstream influence are sharply

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closely with the slipstream boundary,

c) the lift distribution outside of the slipstream shows a

marked increase relative to the corresponding case without slipstream especially for high angles of attack. This

increase extends up to the tunnel walls. where boundary

layer renders the readings uncertain. lts fairly flat distribution and its magnitude suggest a uniform flow modification delaying stalling of the whole wing.

Figure Ua illustrates the lift characteristic of the wing at

its midspan station in uniform flow at various Reynolds Number. In Figure 11b the characteristic of the wing at a point outside of the

slipstream (station 16) in the flow without and with slipstream is shown.

The latter curve does not follow the progressive stalling behaviour of the

first curve.

Figure 12 illustrates the shifting of the center of pressure

with angle of attack at stations inside and outside the propeller

diameter for uniform flow and for flow with slipstream. The difference

in position of the center of pressure leads to the following explanation

of the lift increase outside of the slipstream: progressive trailing-edge

flow separation takes place under uniform flow conditions at angles of

attack

>

70; this separation is reduced in the flow with slipstream, the

most probable cause being boundary layer control. The shift in position

of the center of pressure can also be observed inside of the propeller

diameter. which leads one to believe in a boundary layer modification

caused here by the decrease of the parameter Ap/q due to added

velocity inside the slipstream.

In order to evaluate the overall slipstream effect the total

lift increase AL was rende red nondimensional in the following ways:

bL

a) as T and plotted vs oL in Figure 13a. We note the

increase in ~L with eX. at a rate faster than proportionality and

T

also the non -linearity with

cr ,

b) as

AL

vs

C5

in Figure 13b. This quantity can be termed

slipstream-f~Jng

effectiveness of the wing. The value unity

signifies that ~ L has arisen from a turning of the thrust vector through

an angle 0<.. Curves for different eX. do not coincide,

From a) and b) it results that wUh the exception of

U

near

unity (slipstream only: free jet), turning of the thrust vector does not give a satisfactory evaluation of the phenomenon,

c) as C

6;_

vs

û'

in Figure 14. Here CT

~T'

is the

LW ~q -w

local CL at the midspan station in uniform flow and S the area of the wing

.

-I

(18)

,

\

immersed in the slipstream. Again the curves for different

ex

do not coincide. Values in this representation are affected by the non-linear varia tion of C

Lw

with 0( •

,

dl·

as 2 'Ir0<

t

L

"3:"

vs

cr

in Figure 15. The general shape of the curves remains e sa e, the difference for different ex being somewhat smaller. Comparison with the solution given by the slender body theory» ~Ref. 2), shows good agreement for

cr

near unity, but

higher values for decreasing

cr.

Clearly. in all.these representations, the angle of attack i~

still an independent parameter due to the destalling effect. In order to separate the lift due to boundary layer modification from the totallift increase» we ,de fine :

ALp =

\ ~~

J

(l

with -

l

without slipstream) dy span slipstream "destalled flow"

where (

Z

)fwithout slipstream'was faired in between the portions of the

~tclestalled flow" )

lift distribution curve outside of the slipstream for flow with slipstream

as in the following sketch; this represents an estimate of the lift

distribution for uniform flow with the same boundary layer modification as for flow with slipstream:

t

l

lwith slipstream

ithout slipstream "destalled flow"

'lwithout slipstream

~

~---~~

Thus

~p

is an approximation of the lift increase as it should be predicted by potential flow theory.

•

For comparison ALp is fepresented in the sane non

dimensional forms as ALas shown in Figures 16, 17 and 18. It is se en

that:

ALp

'

.

a)

ex

Tand 27rOCS Aq are Independent of 0( . Further-more toe' values agree very wel! with the solutions given by slender body theory1. The difference between these values

and the

AL

and

DL

curves indicates the magnitude

0< T 2ro< S ~

1The idea of subtraCting the de stalling lift and correlating

the remaining Ifpotential flow" lift increment

6.Lp

with slender body

theory was suggested by Dr. H. S. Ribner.

(19)

of the destalling-eife"Ct. (See Figures 16 and 17).

b) C

D.;P.iq

is not indepem:ient of 0 ( .

(

See

Figure 18).

Lw

VIII. -CGNCLUDING REMARKS

1) Experimental data are presented for compaIlison with and evaluation of theoretical results . The data are given in the form:

spanwise

lift distribution:

for the parameters angie of attack: slipstream strength:

(1jqj:

_<10)

_with and without slipstream 0 0 0

ol

= 6 1/2

11 1/2 16 1/2 vs

0 -

;;

.

~j

- q-o = .26 .49 .62 .76 .88 qj

+

~

No wind tunnel corrections were made to the results .

I ( .

~

fl

Y/R 'I'

Tunnel wall interference (slipstream "image" effects) are presumably small since at small 0(. the added lift due to slipstream falls off

rapidly with the distance from the slipstream axis. Useful comparison

with potential flow theories is expected with the data for 0(. = 6 1/20. Results reflect the characteristics of the particular propeller operating

conditions (see Figures 8 and 9) and are not wholly general. Correction

for angle of rotation should be made in the theoretical treatment. The q profile is not expected to be of significant importance. At higher angies. boundary layer flow plays an important part in the phenomenon.

2) The spanwise lift distribution shows the importance of flow modification outside of the slipstream at ~ngles of attack near stall

in uniform flow. This effect extends as far as

'

y

/R = 4 i. e .• over the fuU wingspan in this experiment. The change of CL and the chordwise aft displacement of the local center of pressure due to slipstream can probably be attributed to boundary làyer modification inside and outside of the slipstream which delays flow sep'aration.

AL

3) Overall quantities s uch as or

ex

T CT ~y. S

m

óL

-w

21l1X S

z:q

do not describe the physical picture meaningfully enough for

6'

=F

1 and for large deflections . However. af ter separation of the de stalling effe cts • the results expressed as

.öL

p .

AL

P conform

0< T 211'0( S

Aq

(20)

.

,

( 13 )

4) Since destalling effe cts are specific to the configuration

considered. general overall parameters governing this part of the

phenomenon are not evident. Among the probable ones are Reynolds

Number. stall pattern, deformation of slipstream. However I in

general, the phenomenon of lift increase due to slipstream is

considered to occur along the lines of the following diagram:

r

f

" .

~r.e~ttoi1.'0f dqwRwash ' ~ Idecrease of

in slipstream

"1

Cliocal

~'

boundary layer

-1

increase of , Lift increase

etabilised in Cl local ~ ~ inside

immersed section

(reduced

pI

q)

slipstream

Slipstream =

/

:Lncreased velocity increased depression above the wing around. immersed't----: ... !1001 increased pressure below the wing

section of the wing boundary layer suction preventing ~ separation outside of ~l1pstream increase of Cl local ~

I .

.

increase of Cr~~.tH~n of up~ash '). C local putside of slipstream '--_..:.1 _ _ _ .... J

Lift inèré'aSE

outsi.de

slipstream

5) From the practical point of view of STOL aircraft, the

main conclusions of this study are:

a) lift gains as predicted by potential flow theories can be

greatly improved in practice by "destalling" effe cts •

b) these effects have to be studies on the specif!c

configuration considered. In this way, the design could

intentionally make use of the slipstream velocity to

(21)

1. 2. 3. 4. 5. Koning, C. Graham, E. W!, Lagerstrom, O. A. , Licher, R. M., Beane, B. J. Rethorst, S. Ribner, H. S. Smelt, R. , Davies, H. 6 . Stuper, J. REFERENCES

INFLUENCE OF THE PROPELLER ON OTHER PARTS OF THE AIRPLANE

STRUCTURE=,., AERODYNAMIC--THEORY,

Edited by W. F. Durand, Vol. IV, Berlin, 1935

•

A PRELIMINARY THEORETICAL INVESTIGA-TION OF THE EFFECTS OF PROPELLER SLIPSTREAM ON WING LIFT, Rep. No. SM-14991, Douglas Aircraft Co., Santa Monica Div., November, 1953

LIFT ON A WING IN A PROPELLER SLIpSTREAM AS RELATED TO LOW SPEED FLIGHT, Eng. Rev., October,

1956, pp. 42-48

THEORY OF WINGS IN SLIPSTREAM$, Unpublished Report, DeHavilland Aircraft of Canada Ltd., March, 1957

,

ESTIMATION OF INCREASE IN LIFT DUE TO SLIPSTREAM

British ARC, R & M 1788, 1937

EINFLUSS DES SCHRAUBENSTRAHLS AUF FLUGEL UND LEITWERK,

Luftfahrtforschung Vol. 15, No. 4, 193Q;

(22)

•

Susp.nsion \--~--- ---S.HlinQ Chomblr Wir . . ~A Sflain gog. L.ods ' - - -.... ~A

Tunnel IongitudinQI section

FIGURE 1

.

_•

_..

..T!!!. Sectio"

A-A

Cross section ot the pivot axis

(23)

.!!. ac

o

..

_o Q. (/) Leods FIGURE 2

VLiEGTUIGE:OU 'UNDE

K2Inaal~tr at 10 - D LFT

(24)

FIGURE 3

(25)

i)

~;,;,. ...

FIGURE 4

(26)

FIGURE 5

EXTERNAL VIEW OF THE FORCE MEASUREMENT APPARATUS (with switch box and straingage indicator in foreground)

(27)

(28)

_ . . . _ _ .. _ _ . _ ... tllW ... ...-JI _ _ _ _ _ _ _ _ ..

.

6~----~---~----~1---~----~----~

105 psi

gl

dia. a"pitch Propeller

/

Q:

.

::I:

·4~----~~----4---~~--~-r---+---~ "-Q)

-Q)

c-o

à:

·3

~

c

o

"-~.2---+---4-~~----~--~--+---~

~

.1

~---+---A---~~----~4---~~----~

o

~~~~

______

~

____

~~

____

~

__

~~~

____

~

o

4000

8000

12000

16000 20000 24000

R.P.M.

FIGURE 7

(29)

,...,

0 (,) ~ (.!) Q)

2

c ~ ~ .r:.

-W C 1..-' cr

°1~.5~~~1~.0~---.~5---0b---~.5~--~1~.0~~~1.5·

20

r - - - ; - - - , - - -

rR

I

~15 ~---+---~---~~~~--+-~~----~ ~ Q)

~IO ~----_1---_+---_H~~~~~~~~----~

c o

-

c

5 -

e

'0

0

.,g? Cl

g

5 <:!:t

10

15

FIGURE 8

SLIPSTREAM CROSS SECTIONS

DYNAMIC PRESSURE AND ROTATION ANGLE (measured 6" downstream of propeller disk)

(30)

I

4 '.c'

L...I

I~

0::

3 t=

-c

c::

c

~2

-

en ::::J ~

~

.2

.4

.6

cr-

=

ëfj

-q,

<fJ

+Q.

.8

1.0

. - . . . : : - - - + - - - - " " " t " " - - - r - - - - . . . , . . . - - - ,

'..,

::K.

8 J---...p....,..:---+---+---+--~

.0

-

6

~---+---+_~~--4---~---~

.

o

ct: ~

A

~---4---~---+--~~~---~

-

_g

~

.2

~---+---+_---4---~--~~~

::>

0

0 .2

.8

FIGURE 9

PROPELLER THRUST SLIPSTREAM DYNAMIC PRESSURE FORCE AND VELOCITY RATIO (varia tion during the experiment. in function of

the slipstream strength parameter)

(31)

1.2

1.0 .8

. 6

.4

. 2

0 u

1.0 ~.8

,...:-...

_.6

11 <o..J

.4

. 2

0 .8

.6

.4

.2

0 .

9;r-Qo

!

~

0-

=

ëf.+Qo

=.26

It _1\

L

~

.03

V

I

I~

-~ ~

It

~

-

.... Ä <ti. ....

...

""'" I~

-4

I with

o

slipstream . wÏ1flouf

<>

slipstream. I / ' ~

/

'"

~

"

-

I

~

7' ~

<>

...

o

r

- Q

<>'

IV

-

....,

p

I I

),~

,.co; I ~ -. ~ ~ ... \.Jo ~ ..-

p...r

_

...

-.~

"V . " I

3

2 o

_, / I- b

~R

FIGURE lOa

SpANWISE LIFT DISTRIBUTION

<:5 ::

.26 oe.

=16,0

~~

'"

<>~

~

\

0(.

=lIt 0

~o

~

...

0(.=6,0

... ~ ~

~o

~

(32)

-o

1.2

1.0 .8

.6

.4

.2

o

1.0 -

.8

lIN

IJl

s:

.6

.4

11

<u-1

.2

o

.8

.6

.4

2 o

4

(J=

qj-

qo

=~9

1.

~

7,

0\

qj+qo

7

1 0 , . / ~ ~

h

7;

<> <> ~ <;> .". _<> -.;;;;00' ... <> ~

1

I

,

o

with slipstream _L-. without <> slipstream

(,

,

7 ~

~

0 _17-_1000.. 0<>

p7

_I

\

~ <;> 'V' <> -v I

-r

_,

I

~f\

_,

I

~

1"1"- ~ -, A ...

"'

-r

--

_fV<>

1<>

<>

0

r

-.;;, I

I

3

2 o

%

FIGURE lOb

SPANWISE LIFT DISTRIBUTION

cs

= .49 0

~

=16!-"

v

~""t;

...,.

_...

~. 0

CX=1I1-2 ...

~

ot.

= 61

! 2

-

_~

2

3

4

(33)

o

I.

2

1.0 .8

.6

.4

.2

o

0 1'1.

._ (\J

10-\ : ' .8

~

S

.4

2 o

.8

.6

.4

2 o

4

0"=

qj-qo=

.62

lij

+qo

_~=

16t!

~ I

\

I I

,

9'

0

""""

~ ""0.. 0

I}

~

0 _'\.. _.L1"

,

r

O.g

e>_ 0"""" r- ~ - 0

~

I

o

with slipstream .-<> without

t,

slipstream

I

(

I

_~

-

j

.-cf'

v -~

_~

_<;

₀

-

"

I ~

/,

,

0

j

~

é) _~0 x '""'" ...: In I ~

-~

~

-I

3

2 o

%

FIGURE lOc

cr'

=

.62

Ol= Ilf!

--ex=

6t!

.- .-n ~

-,..,

2

3

4

(34)

I.

2

I.

0

8

6 .4

2 o

o

---..

~

I

.0

+C\J

lëT

~.

8

6

11

<u...J

A

2 o

8

6 .4

2 o

4

I

cr-.

qj -

q,=

.

76

0

ql+qo

sr-c\

0!=16~-/

I

/

I - ... .N

~

~-n.

o/:

0 0- <>

_~-

_~

~

- -

~ ~I

--~

Ir

I _<>~ I 0 with slipstream without 0

<> _{• I,atnam}

r

_1('\

_eX=

lIi-/

/1

0 _1"\ <>

l<>

0

~

ó J')o" _oL> <> ...

_,

-v _:""'10.. _... <> <> 0--

-cr

0 , I

/"'Ol..

!

,

J

\

/

~

,... ... <> ~

-... 0.

-

,. _J _<>

-

_~ ~

--

-

~-c>

--

<>

3

2 o

~R

FIGURE lOd

cr

=

.76 0 0 ~=

6k-"

:9-...

~

2

3

4

(35)

1.0 .8

.6

:4

.2

0

1.0 .8

(J

~

.6

I'

~

:4

11

_.2

<u..J

0 .8

.6

.4

.2

0 -.2

0-

=

Qj

-qo=

.88

Ii\

I

\

ëfj

+qo

_Ol

₌

₁₆₁

_! 2 I

~

I

~

,

\

---

~ (S

\.-

~ 0 ~ 0

_o

_:-0

0

-

0 0 0 0 0 0 I with

I

C\

0 - 0 -_ar _Ipstream

_I

Ol

=

11.1.

-2 extropolated

\

- - for flow without

slipstream I 0

)

~

hS"-O-

.~ _~ I

ç

Po--~

0 0 0 t'\ 0

~

Ol=

61

g 2

1 _\

~

I

-..:::;

--

._~ 0

--

o ....

I\.

.L'

-

_.... 0 0 0 _0"

-0 0

\

I

\j

4

3

2 ₂

3

4

FIGURE lOe

(36)

1.6

1.4 ~=

₍

l

)mld'pa~

1.2

1.0

~

.8

u'"

.6

.4

.2

o

0 ₅

₁₀

₁₅

₂₀

0(

[Oeg.]

1.6r---r---r---,

1.4 C

_Lw

=

(l

).tation 16 / 16

~~c:

1.2

station 16 I

VR

= 2.2!S

with

--asli

pstrea m

10

. ~---+---~~~---~~Wl

· th

out

I

:.8

slipstream

u.J

.6

~---~#---~---4_---~

A

~--~~~---4---+_---~

.2

~~----~---~---~---_4

0

5

10

15

20 ex...

[Oeg.]

FIGURE 11

a) LOCAL LIFT COEFFICIENTS IN UNIFORM FLOW AT MIDSPAN

b) LOCAL LIFT COEFFICIENTS IN FLOW WITH AND WITHOUT SLIPSTREAM AT A SECTION OUTSIDE OF THE SLIPSTREAM

(37)

f-15

Station 16.

10

1 - \ - \ - - - - + - - - + (outside of ) slipstream

5 --0-,...,

0

0' Q) ,"'C •

Station

10

~ (,) 0

--

0

...

0 Q) 0'

c

0 f t

~

10

5

0

15

10

5

0

t-+~r----+---+ ( i nside of ) 1 - - - 0 , - \ - - - - + - - - + sHpstreom

0 X

_cp t

-

_---

-0...

Station

4

( outs id e of ) ~.---t---+ slipstream

----

..

-2

[in]

o

2 [in]

distance of center of pressure aft of

~

chord point.

with slipstream

- - - without slipstream

Resu \tont Lift sJipstrea m

Resultant Lift

shlfted aft

sllpstreom TE

Partia Ify seporated flow

lE

11 Destol1 e d" flow

TE

-FIGURE 12

VARIA TION OF THE POSITION OF THE CENTER OF PRESSURE WITH

ANGLE OF ATTACK, FOR FLOW WITH AND WITHOUT SLIPSTREAM, AT DIFFERENT SPANWISE STATIONS