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NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

TECHNICAL NOTE 3918

WIND-TUNNEL INVESTIGATION OF EFFECT

OF PROPELLER SLIPSTREAMS ON AERODYNAMIC CHARACTERISTICS

OF A WING EQUTPPED WITH A 50-PERCENT-CHORD SLIDING FLAP

AND A 30-PERCENT-CHORD SLOTTED FLAP

By Richard E. Kuhn and William C. Hayes, Jr.

Langley Aeronautical Laboratory Langley Field, Va.

L LlBWRwr copgr Washington

Feb mazy 19 57 FEB 26 1957 LANGLEYAEBOHALITICAL L);BO&JTORy

U&R&RI, NACA LANGLEY I-~Y: y vlHCilNIA

NATIONALADVISORYCOMMITTEEFORAERONAUTICS

TECBRICAL NOTE 3918

WIRD-TUNNEL INVESTIGATION OF EFFECT

OF PROPELLER SLIPSTREAMS ON AERODYNAMIC CHARACTERISTICS

OF A WING EQUIPPED WITH A 50-PERCENT-CHORD SLIDING FLAP

AND A 30-PERCENT-CHORD SLOTTED FLAP

By Richard E. Kuhn and William C. Hayes, Jr.

SUMMARY

An investigation of the aerodynamic characteristics of a wFng equipped with a 50-percent-chord sliding flap and a 30-percent-chord slotted flap operating in the slipstreams of two large-diameter propellers has been conducted in the Langley 300 MPH 7- by lo-foot tunnel.

Large tunnel-wall effects for which there are no known correction methods were encountered in the tests. However, because of the current interest in and general scarcity of data applicable to aircraft designed for vertical take-off and landing (VTOL) and for short take-off and landing (STOL), the results (uncorrected) are presented herein with only limited discussion. It was observed, however, that stalling would occur in conditions approaching steady level flight at high-power conditions, but that a leading-edge slat effectively delayed this stall.

INTRODUCTION

An investigation of the aerodynamic characteristics of wing- propeller combinations that may be applicable to aircraft designed for vertical take-off and landing (VTOL) and short take-off and landing (STOL) is beFn@; conducted in the Langley 300 MPH 7- by lo- foot tunnel. The aerodynamic characteristics at low forward speed of a wing-propeller combination without flaps (tilting-wing configuration) at angles of attack up to 90' is reported in reference 1. Refer- ences 2 and 3 report the characteristics of this same model equipped with plain and slotted flaps, respectively.

2 NACA TN 3918

In the present-investigation the aerodynamic characteristics of a semispan-wing model equipped with a 50-percent-chord sliding flap, a 30-percent-chord slotted flap, a 30-percent-chord leading-edge slat, and two large-diameter propellers have been investigated. The characteristics of this model at zero forward speed are reported in reference 4.

c

SYMBOLS -..

When a wing is located in the slipstream of a propeller, large forces and moments can be produced even though the Tree-stream velocity decreases to zero. For this condition, coefficients based on the free- stream dynamic pressure approach infinity and become meaningless. It appears appropriate+herefore, to base the coefficients on the dynamic pressure in the slipstream. The coefficients based on this dynamic pressure are indicated in the present paper by the use of a double prime. The relations between the thrust and the dynamic pressure and velocity-- in the slipstream have been derived in reference 1. The positive sense of forces, moments, and angles is indicated in figure 1. The pitching moments are presented with reference to the center of gravity shown in ' figure 2.

-. a

B

b

b*75R

number of propeller blades

twice span of semispan wing, ft

propeller blade width at??-percent-radius station, ft

cL lift coefficient,' Lift m

CL"

Cm

lift coefficient, Lift -2 q"S/2

pitching-moment My coefficient, - qEs/2

Cm" pitching-moment

C m,p"

cN,p"

pitching-moment

normal-force coefficient of propeller, q"s/2

coefficient, My q"ES/2

coefficient of propeller, My,P q"%/2

.N, c

-

*

NACA TN 3918 3

c

.

cP

CT

CX

Cx"

Tcl'

T " c,nom

C

E

cS

D

L

MY

%P

NP

n

&

9

power coefficient, !Z!E!&- pn3D5

T thrust coefficient, - 2D4 pn

FX longitudinal-force coefficient, - G/2

FX longitudinal-force coefficient, - q"s/2

thrust coefficient, Z q"P2

nominal value of thrust coefficient (taken as the average value at low angles of attack)

wing chord, f-t

mean aerodynamic chord, 2 s b/2

c2dy, ft 0

slat chord, ft

propeller diameter, ft

lift, lb.

pitching moment, ft-lb

propeller pitching moment, ft-lb

propeller normal force, lb

propeller rotational speed, rps

torque, ft-lb

free-stream dynamic pressure, i oV2, lb/sq ft

4 .-- NACA TN 3918

q” T

S

T

V

FX

a

‘-75R

6

P

slipstream dynamic pressure, q + -, f D2

lb/sq ft

twice area of semispan wing, sq ft-

thrust per propeller, lb

free-stream velocity, ft/sec

longitudinal force, lb

angle of attack, deg

propeller blade angle atT5-percent-radius station, deg

surface deflection, deg

a

mass density of air, slugs/cu ft

propeller solidity, z .

Subscripts:

50 50-percent-chord‘ sliding flap

30 TO-percent-chord slotted flap

i inboard propeller.

0 outboard propeller

S slat

h horizontal tail

f flap

-

MODELANDAPPARATUS

A semispan-wing model-of a hypothetical four-propeller airplane was used in the present investigation. A drawing of the model with perti- nent dimensions is presented as figure 2, and a phot0grap.h of the model- mounted for testing is shown in figure 3. The principal geometric characteristics of the model are given in the following table:

NACA TN 3918 5

Wing: Area(semispan), sqft .................... Semispan,ft ......................... Mean aerodynamic chord, ft .................. 1.51 Aspect ratio ......................... 4.89 Taper ratio ......................... 0.80 Airfoil section ..................... NACA 4415

Horizontal tail: Area (semispan), sq ft .................... 1.65 Semispan, ft. ........................ 2.0 Mean aerodynamic chord, ft .................. 0.83 Aspect ratio. ........................ 4.85 Taperratio ......................... 0.66 Airfoil section .. . . . .. ... Tail length (from'E/& of w&g'& E/4 of tail), ft ...........

NACA 0012 . 3.76

Propellers: Diameter,ft ......................... 2.0 Airfoil section ...................... ClarkY Solidity (per propeller), a ................. 0.07 Number ofblades. ...................... 3

The ordinates of the slotted flap were derived.from the slotted flap 2-h of reference .5 and are presented in table I along with a sketch of the profile of the sliding flap. The cross section of the leading- edge slat is also shown in table I. For these tests the upper surface of the wing was not modified as it would have to be in a practical appli- cation in order to retract the slat; however, it is believed that this difference would have only a SIXKU effect on the results. The end plate, which was installed for all these tests, was made of l/16-inch aluminum and is shown in figure 4. It was located 10.3 inches outboard of the center line of the outboard propeller.

The propellers were driven by variable-frequency electric motors from a common power supply. The speed of each propeller was determined by observing stroboscopic-type indicators to which were fed the output frequencies of small alternators connected to each motor shaft. Because the propeller blade angles were adjusted to give the same thrust from both propellers at zero angle of attack, their rotational speeds were usually matched within 10 r-pm. The outboard propeller rotated against the tip vortex (right-hand rotation on right wing as tested) and the inboard propeller rotated in the opposite direction,

The motors were mounted inside nacelles through strain-gage beams (as shown in ref. 1) so that the thrust and torque of each propeller could be measured. The inboard nacelle was equipped with additional

6 NACA TN 3918

. instrumentation so that the propeller norm&l force .&nd pitching moment could also be meatiured. The total model lift; longitudinal force, and pitching moment were measured on a balance at the root of the u3ng.

TFSTS

The investigation was conducted in the Langley 300 MPH 7- by lo- foot tunnel. The tests were made at various free-stream dynamic pres- sures and propeller thrusts so selected as to maintain a dynamic pressure of about 4.8 pounds per square foot-in the slipstream. The tests with the propellers off were also run at a dynsmic pressure of 4.8. The desired thrust was obtained on both propellers at zero angle of attack by appropriate adjustment in the propeller blade an&es and the thrust of the inboard propeller was held constant through&t the angle-of- attack range. Because of different inflow conditions, the thrust of the outboard propeller varied slightly from the desire& value at angles of attack other than zero.

The Reynolds number in the sli sp

stream, based on the mean aerodynamic chord of 1.505 feet?as 0.62 x 10 .

In order to minimize the time required for #the tests, the operating conditions were chosen so that only two were required. A blade angle of about 79

repeller blade-angle settlngs was used for tests at-the

higher thrust coefficients, and a blade angle of about 20' was used for the lower thrust coefficients.

t.’

TUNN-EL-WAIL EFFECTS AND CORFGCTIONS

Large effects of the tunnel walls on the data were encountered during the tests. These effects are shown in figure 5 where the data obtained at zero forward speed in the tunnel are compared with the results obtained from tests in a large room (ref. 4). The tests in the tunnel were made with a curtain suspended in the diffuser of the tunnel so as to prevent the propellers of the model-from setting up a recircula- tion of air in the tunnel; In addition, the doors into the tunnel (imme- diately downstream of the test section) were open to prev-enta circula- tion from being set-up within the qes+secGon.

The efiecti-of &he tunnel-wall restrictions have been determined only for the case of zero forward speed (Tc" = 1.0) and for the flap deflections shown in figure 5, but they are probably also present to an unknown extent at other thrust coefficienmand at lower flap deflections (particularly at high angles of attack). Procedures for correcting for these effects are nut know-f, . .

c

NACA TN 3918 7

The power-off data and the data for low angles of attack and low flap deflections (power-on) are probably not appreciably influenced by these effects. Normal potential-flow corrections for the effect of the constraint of the tunnel walls and the blockage of the model have been applied by the methods of references 1, 6, and 7. In determining these corrections only the aerodynamic forces were used. (Direct propeller thrust was subtracted from the data.)

PRESENTATION OF RBULTS

The results of the investigation are presented in the following figures:

Figures

Tunnel-wall effects at ICC" = 1.0 . . . . . . . . . . . . . . 5

Data with propellers and nacelles off: # Effect of flap deflection (slat off) . . . . . . . . . . . 6 to 8

Effectofslat...................... 9

Effect of slipstream: Flaps neutral ...................... 10 Flaps deflected ..................... llandl.2 Effect of slat ...................... 12 and 13 Effect of stabilizer ................... 14 to 16

I Propeller characteristics . . . . . . : . . . . . . . . . . . 17 and 18

DISCUSSION

Because of the current interest in flight at very low speeds and the general scarcity of data of the type obtained in this investigation, the test results are presented herein, although it is tiown that the data include large tunnel-wall effects. No detailed discussion is presented; however, two results of the investigation should be noted.

Although the wing is almost completely immersed in the slipstreams, the data indicate that, with the slat off, stalling would be encountered in steady level flight or in conditions approaching steady level flight. This result is indicated for the wing with the flaps neutral in figure 10(e). on this plot CxH = 0 indicates steady level flight, negative values of CXlr indicate decelerating or gliding flight, and positive values of Cxn indicate accelerating or climbing flight. For the

8’ NACA TN 3918

three highest thrust-coefficient condittons tested, maximum lift was encountered at positive values of CX". At CX" r-0 (steady level flight), the drop in lift indicates that the wing i_s stalled.

With the rear flap deflected (fig. 11(e)) the stall is delayed somewhat but is more severe. With both flaps deflected (fig. 12(e)) the data indicate that level flight could be achieved without-the wing stalling; however, in order to make a landing approach, x" must be negative and under these conditions the wing would stall violently.

Adding the leading-edge slat (fig. 13(e)) -e-ffectively delays the stall. No attempt-was made to determine an optimum--slat configuration for stall control. The slat arrangement used was chosen for good pitching-moment and ground-effect characteristics in hovering flight.

An examination of the pitching-moment data at low forward speeds (figs. 12, 13, 15, and 16) indicates tha-t;; with the incidence settings used herein, the flow on the horizontal tail is stalled on the lower surface. The pitching-moment data with the horizontal tail on, therefore, are not representative of what could probably be obtainedwith a complete conf'iguration, but-are included because they are believed-tobe of general interest and because the slat-effectiveness data are also shown in these figures.

.

.

CONCLUDING REMARKS

Stalling was observed to occur on a wing equipped with a 50-percent- . chord sliding flap and a 30-percent-chord slotted Map operating in the slipstream of two large-diameter propellers in conditions approaching steady level flight-at high-power conditions; however, this stall was . effectively delayed by the addition of a leading-edge slat.

Large tunnel-wall effects for which there are no known corrections were encountered.

Langley Aeronautical Laboratory, National Advisory Committee for Aeronautics, -

Langley Field, Va., October 23, 1956.

2F NACA TN 3918

REFERENCES

Y

1. Kuhn, Richard E., and Draper, John W.: Investigation of the Aero- dynamic Characteristics of a Model Wing-Propeller Combination and of the Wing and Propeller Separately at Angles of Attack Up to 90'. NACA Rep. 1263, 1956. (Supersedes NACA TN 3304 by Draper and Kuhn.)

2. Kuhn, Richard E., and Draper, John W.: An Investigation of a Wing- Propeller Configuration Rnploying Large-Chord Plain Flaps and Large-Diameter Propellers for Low-Speed Flight and Vertical Take- Off. NACA TN 3307, 1954.

3. Kuhn, Richard E., and Draper, John W.: Investigation of Effectiveness of Large-Chord Slotted Flaps in Deflecting Propeller Slipstreams Downward for Vertical Take-Off and Low-Speed Flight. NACA TN 3364, 1955..

4. Kuhn, Richard E.: Investigation of Effectiveness of a Wing Equipped With a 50-Percent-Chord Sliding Flap, a 30-Percent-Chord Slotted Flap, and a 30-Percent-Chord Slat in Deflecting Propeller Slip- streams Downward for Vertical Take-Off. NAGA TN 3919, 1957.

5. Wenzinger, Carl J., and Harris, Thomas A.: Wind-Tunnel Investigation of an N.A.C.A. 23012 Airfoil With Various Arrangements of Slotted Flaps. NACA Rep. 664, 1939.

6. Gillis, Clarence L., Polhsmus, Edward C., and Gray, Joseph L., Jr.: Charts for Determining Jet-Boundary Corrections for Complete Models in 7- by lo-Foot Closed Rectangular Wind Tunnels. NACA WR L-123, 1945. (Formerly NACA ARR L5G31.)

7. Herriot, John G.: Blockage Corrections for Three-Dimensional-Flow Closed-Throat Wind Tunnels, With Consideration of the Effect of Compressibility. NACA Rep. 995, 1930. (Supersedes NACA RM ~71~28.)

10 NACA TN 3918

Slotted-flap ordinates

station, fraction of wFDg chor&

"12

.67

.68

.@

.70 1z

.73

.74

.75

.76

.77

.78

:iZ .81

:g

Ordinate, fraction of ring chord

nap nose Slot .

Upper Lower

-0.0183 ----me -m-m-- -.0175 ------ ------ -.ol6g ------ ------ -.0145 ----mm ----mm .-.00&I ------ ------

.Olh 0.0075 0.0075 -----mm .0375

.02& .02&I .0210 .03w ::gg -.0020

.0430

.0470 :%; -.oll5

-.Ol2!5 .ow .0490 -.ol25 .0520 .0510 ---*mm .05% .ogm ------ :g;; - 0525 ------

.0530 .OP5 .05= ------ ----me

.0531 .0515 ----mm

------- ------- .we .0485 ------ ----mm

St8tti, fractson of slat chord

0 .025 .W .075 .loo .150

:g

:E .5=J .&cl

:'Z .900

l.ooo

slat oralnatcs I Ordinatea, fraction of

slat chord

Upper

0 .060 .083 .w7 .1op .u5 .lm .I22

2% .m5 .WJ .w3 2; .013

1 Lower

0 -.og -.027 -.025 -.023 -.cQ7 .Ol2 .os .046 .070 .0%3 .071 .056 .039 .020

0

Figure l.- Conventions used to define positive sense of forces, moments, and angles.

Model cbamckrktics

Wing: Arecrhwnk&sqft 5.50 Aspect ratio 489 Taper ratio 0.80 section NACA 4415

Horizontal toi/: Area &mkpanA sq ft I.65 Aspect ratlo 485 Taper ratio 0.661 Sect/on NACA 0012

Propellers:

Figure 2.- Drawing of model. All dbmx3ions are in inches.

r I

I , L I I I

L-91043 Figure J.- Photograph of the model installed on the ceiling of the Langley 300 KPH 7- by lo-foot

tunnel. u = 40'. G

14

Flaps neutraj

Flaps def tected

Figure 4.- Details of the end plate.

NACA TN 3918

-2

.4

.6

.8

ID/ IO

.5

CL” 0

Cm "0

. Figure 5*- Effect of tunnel-wall constraint on aerodynamic charac

tics at zero forward speed.

15

teris-

16 NACA TN 3918

40 -20 0 20 40 60

(a) Lift--coefNcient.

Figure 6.- Effect of-flap deflection. 6f,w = O"; propellers a off; slat off; stabilizer off.

,nd nacelles c

3F NACA TN 3918

-.6

0 o- n 20

40

-4

-2

(b) Longitudinal-force coefficient.

Figure 6.- Continued.

18 NACA TN 3918

sr; 309 e7

0 0 020 0 40

-4 ~““‘r”‘~““‘i”‘r”““‘~““““‘~““r~“‘r~”f -40 -20 0 20 40 60

a, m7 (c) Pitching-moment coefficient.


.

NACA TN 3918

3.0

2.5

2.0

/.5

I.0

.5

0

0 -2 -4 -.6 ~8 -10 CX

(d) Variation of CL with %.

Figure 6.- Concluded.

20

deg 0 0 q ?O

NAC!A TN 3918 -

.

-60 -40 -20 0 20 40

aJdeg

(a) Lift coefficient. c

Figure 7.- Effect of flap deflection. 6f,w = 50’; slat=off; stabilizer

off; propellers and nacelles off.

NACA TN 3918

-60 -40 -20

(b) Imgitudinal-force coefficient.


22 NACA TN 3918

.

8f, 309

Qe7 0 0 q 20

(c) Pitching-moment coefficients .


NACA TN 3918 23

.

3D

25

20

L5

/.O

.5

0

-. 5 .2 0 -2 -4 -.6 -.8 40 -12 -/.4 - 16

CX

(d) Variation of 'I& with $.


24 NACA TN 391-8

.

.

-.5 I.! !.! !.J ?L’ ’ ’ ’ ’ ” “‘1. ” ’

-140 -20 0 20 40

a, de7

60

(a) Lift coefficient.

Figure 8.- Effect of flap deflection. Sf,30 = 0'; propellers -.

off; slat-off; stabilizer off.

.

and nacelles

4F NACA TN 3918 25

. CX

Sf ,501

deq 0 0 n 30

42 0 50

40

-.8

-.6



26 NACA TN 3918

r

&fJ 50’ w

0 0 0 30 0 50 _

.3

.2

./

0

-.I

#2 -40 -20 0 20 4c

(c) Pitching-moment coefficient.


3

.

NACA TN 3918

Sf850,

m? Q 0 q 30 0 50

(d) Variation of c tith (3~.

2


NACA TN 3918 28

0 Slat off

0 Slaf on


Figure 9.- Effect of slat. Ef 5. = 50'; 6f,30 =. 40'; stabilizer off;

propelleks and nacelles off.

.

NACA TN 391-8

0 S/of off q S/of on

-60 -40 -20 0 20 40

QJ e7



30 NACA TN 3918

0 S/of off

0 S/of on 4

.3

2

./

0

4

5

L

-60 -40 -20 0 20 40 Q, deg



NACA TN 3918 31

0 Slat off

q S/of on

.Y

0 -2 -4 -.6 - .8 -1.0 -/.2 -14 -Id6 4.8

CX

(d) Variation of CL with Cx.


32 NACA TN 3918

2.0

I.5

I.0

CL

5

0

-LO

T c, noi

q 0 0 24 A .49 b .69 b .90

-40 -20 0 20 40 60 80 /oo /. cr, deg

(a) Lift-coefficient.

Figure lO.- ETfect of slipstream. +5o = O"; 6f,30 = 0”; slat off; stabilizer off.

5F NACA '33 3918 33

.

0 0 0 24 A 49 b 69 P .90


Figure lo.- Continued.

34 NACA TN 3918

I q 0 0 .24 A .49 b .69 h .90

.

-40 -20 0 20 40 60

Q, dw



NACA TN 391.8 35

.8

.6

T,” .4

-40 -20 0 20 40 60 623 /Gy) 120

a., deg

(d) Thrust coefficient.


T lr c, nom

q 0 0 .24 A .49 b .69 b .90

.

G,nom Y

0 0 0 .24 A .49 a.69 nxl

-42 ID .I9 4 P 0 .P

C,’

(e) Variation of CJ," with CX".

Figure lo.- Concluded.

I .

NACA T?X 3918 37

q .20 0 .48 A .68 b .89 b .94

-60 -40 -20 0 20 40 60 80 x30 /20


Figure XL.- Effect of slipstream. %w = 00; tif30 = 4o”; sla.t off;

stabilizer off.

38 NACA TN 391.8

0 .20 0 .48

-60 -40 -20 0 20 40 60 80 fO0 I20 =, dep


Figure ll.- Continued. -

.

NACA TN 397-8 39

jr /I c,nom

q .20 0 .48 A .68 b .89 h .94

L60 -40 -20 0 20 40 60 80 A90 120

a, de7


Figure lL- Continued.

4-o

+!I P

NACA TN 3918

c

r /I c,nom

q .20 0 .48 A .68 h .89 b .94

2-O -40 -20 0 .a9 40 60 80 100 120 ff, deg

.


Figure ll.- Continued.

c

0 .20 o .48 A .68 a.89 b -94

CL”

..- 12 1.0 .8 .6 4 2 0 32 -4 -.6 ~8 -/Xl -I2 -M -Ati

(e) Variation of CL' with CX'.

Figure ll.- Concluded.

42 NACA TN 3918

T # c, nom

0 -48 P .67 b .88 h -92

-80 -60 -40 -PO 0 20 40 60 80 /oo /20 0, o’w


Figure l2.- Effect of slipstream. b,s = 50°; fjf,% = LOO; slat off; & = 100.

c

.

NACA TN 35~8 43

7 /I c, nom

0 .48 A .67 fL .88 b .92

cx”

‘-yBO -60 -40 -20 0 20 40 60 80 KU /20 Q, &7

(b) Longitudinal-force coeffiaent.

Figure l2.- Continued.

44 NACATN 3918

Cm

-.6 -80 -60 -40 -20 0 20 40 60 80 IO0 /20

a,dW

0 .48 A .67 b -88 n .92

(c) Pitching-moment-coefficient.

Figure l2.- Continued.

,

0 .48 A .67 b .88 0 32

-60 -40 -20 0 20 40 60 80 #XI I.50

(cl) !lltirwt coefficient.


46 NACA TN 391.8

0 .48 A .67 b .88

.8 .6 0 12 .-A -.6 -.8 4.0 -12 Cx"

(e) Variation of CL' with Cx".

,

Figure l2.- Concluded.

NACA TN 3918 47

0 .47 A .66 h .86 b .92

..- -80 -60 -40 -20 0 20 40 60 80 LC’ 120

Q, deg


Figure 13.- Effect of slipstream. Q,y = 500; 8f,30 = 400; 6, = 00; sh = 100.

.

48 NACA TN 391.8

”

.6

.8

- -80 -60 -40 -20 0 20 40 60 80 100 12 a, deg

-I.4

-62

4.0

-6

-.R

52

0

.

0 .47 A 66 b -86 b .92


.

Figure 13.- Continued. .

NACA TN 3918 49

0 .47 A .66 b .86 h .92

0

~6 -80 -80 -40 -20 0 SC 40 60 80 IO0 120

a; dw


Figure 13. - Continued.

50 NACA TN 3918

.

.

G, nom N

0 .47 A .66 h .86 b .92

r80 -60 -40 -20 0 20 40 60 80 KW 120

6 W

.


Figure 13 .- Continmd.

. . I

G, nom n

0 .47 4 46 b .86 h .92

.8 .6 .2 0 -.2 -4 -.6 Is 4.0 -/2 G"

(e) Variation of C$," with $'.

Figure Il3.- c0nc1uaea.

52 NACA TN 391.8

. 1 ‘I c, nom

4 .47 A .66 h .86 h .so

2.5

20

I.5

Cj/

/O

4

0

-.5

-80 -60 -40 -20 0 20 40 60 80 /oO Et7 Q, dw


Figure lb.- Effect of slipstream. SfJ y = 5o”; 6f,JO = 4o”; 6, = loo; stabilizer off.

NACA TN 3918

-.8

.6

.8

53

G, tnm II

0 .47 A .66 b .86 h -90

-80 -60 -40 -20 0 20 40 60 80 /oO 120

a, dw



54 -, NACA TN 3918

r /’ C, nom

0 .47 A .66 b -86 h .so

.

cm

‘180 -60 -40 -20 0 20 40 60 80 /LX' /20 0; dw


Figure lb.- Continued.

4 9 r ” C,fJOl?l

0 .47 A -66 h .86 b 30

‘I80 -60 -40 -20 0 20 40 60 80 I&' I20 a'r dw

(a) !bUst‘coeffYcient.

Figure 14.- Contbmed.

56 MACA TN 3918

0 .47 A -66 b .86 h .so

.8 .6 .2 0 72 -4 -.6 -.8 -LO -L2 CX"

(e) Variation of CL' with $ll.

Figure lb-c Concluded.

8~ NACA TN 3918

e

35

3c

25

2.0

I.5

GM I.0

.5

0

57

Tc, nom /I

0 .46 A .65 b .86 h .s/

..- -80 -60 -40 -20 0 20 40 60 80 100 Et3

a, dw


Figure lg.- Effect of slipstream. 6f,30 = 4o"; 6f, y-j = 500; s, = LOO; 8-h = 00.

58 NACA TN 3918

-LO

-.8

76

-4

-.2

cx” 0

.2

.6

8

0 .46 A .65 h .86

-80 -60 -40 -20 0 20 40 60 80 /oo /m

Q, deg

(b) Longitudinal-force coefficient,

Figure 15 .- Continued. .

NACA TN 3918 59

7- ” cpom

b B6 b -91

.- -80 -60 -40 40 0 20 40 60 80 hW I20

a, de7



60 NACA TN 3918

.8

.6

0 .46 A .6.5 b .86 h .9/

-80 -60 -40 -20 0 20 40 60 80 100 120 a,deg

.

.



NACA TN 3918 61

0 .46 A .65 b ,86 b -91

(e) Variation of Ev with $ll.


62 NACA TN 3918

I.5

CL” I.0

3

0

-5

r H c, nom

0 .46 A .66 b .86 h .90

-80 -60 -40 -20 0 20 40 60 80 Im /20

Q, de7


Figure 16.- Effect of slipstream. +,30 = 40°j 6f,50 cm f30°j S, = 10';

s, = loo.

.

.

NACA TN 3918 63

0 .46 A .66 h .86 b .90

180 -60 -40 -20 0 20 40 60 &9 100 120

0, dw



64 NACA TN 3918

r H c, nom

0 .46 A .66 h .86 h .90

-60 .-40 -20 0 20 40 60 80 100 120 a, deg

(c) Pitching-moment coeffihient.


t

-

NACA TN 3918 65

o .46 A .66 b .86 n .90

-60 -40 -20 0 20 40 60 80 100 I20

a, deg



66 NACA TN 3918

.

r * C, nom

0 .46 A .66 b .86 0 .90

(e) Variation of CLH with CxlI.

.


MACA TN 3918 67

r c, nom’

0 .24 A .49 h .69 b 30

0, dw

(a) Thrust coefficient. Outboard propeller.

Figure 17.- Propeller characteristics. +,p = o"; +,30 = O0 j Slat Offj

stabSLizer off.

NACA TN 3918

T c,nom”

0 .24 A .49 b .69 h

-40 -m 0 20 40 60 80 100 f20 4, e

(b) Thrust coefficient. Inboard propeller.


NACA TN 3918 69

T ‘I c, nom

0 .24 A .49 b .69

-20 0 -20 60 120

(c) Power coefficient. Outboard propeller.


70 NACA TN 3918

T ” c, nom

0 .24 A .49 b .69

” . -40 -20 0 20 40 60 80 100 I20

a; deg

(d) Power coefficient. Inboard propeller.


NACA TN 3918

rc, nom* 0 .24 A .49 b .69

. \

(e) Variation of V/nD with a.


72

cm,P ,,

NACA TN 3918

T C, nom “B

: .24 0 ;z A .49 18

b .69 Es 30 F

-- -20 0 20 40 60 80 IO0 Et?

Q., deg

Figwe 18.- Propeller normal-force and pitching-moment characteristics. Inboard propeller; 6f,30 = O"j +p-= O"j s1a-i; off; stabilizer off.

NACA - Langley Field, “a.

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