SUBSONIC TWO-DIMENSIONAL WIND TUNNEL ;NV

OF THE HIGH LIFT CAPABILITY OF CIRCUI

CONTROL WING SECTIONS

by

Robert J. Enqow

Approved fcr Public Relome:S~~Dizr~ibutian Unlimited •••

or

April1975

, III

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OF•-- THE H IH L IF CAAI•IT OFi CIRI

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23. (continued)

d nozzle pressure ratio. Maximum lift coefficients roughly triple those ofthe flapped conventional sections were generatea at incidence slightly lessthan the conventional stall angles and at blowing rates obtainable from bleedof st~ate-of-the-art turbojet engines. An experimental investigation into the Iilift augmenting effects of pulsed unsteady blowing was conducted on a smallerradius trailing edge configuration. An additional investigation was conducted

to determine the effects of spoilers or similar disturbances ahead of the Jetexit. The results of the above investigations provide a data base for thelation Control trailing edges to increase their STOL capability.

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TABLE OF CONTENTS

[~1 PageSUMM0ARY . . . . . . . . . . . . . .. .. .. . .. . . . .. .. . .

INTRODUCTION ............................ ... 1

DESIGN CONSIDERATIONS .. O. ............ ........ ... 3

MODELS AND EXPERIMENTAL APPARATUS ............. ................. 4

iiMODELS ........................ ........................... 4

EXPERIMENTAL APPARATUS AND TECHNIQUE ............ 5

RESULTS AND DISCUSSION .. ....................... 6

NACt. 66-210/CCW AIRFOIL ............. ... ................. 6

Lift Augmentation ............................... 6

Reynolds Numbe) Effects ......................... 9

Drag ........................................ 9j Pitching Moment .................. ..................... 10

NACA 64A-212iCCW AIRFOIL ................ ....... . . 10

Lift Augmentation ............ ................... . l

Slot Height Variation .......... .................. 13

Reynolds Number Effects ............ ................. 14

Drag ........ ........................... 15

Pitching Moment ...................... 15

PULSED BLOWING INVESTIGATION .............. 16

SPOILER INVESTIGArION ........... ................. ... 19

CONCLUSIONS ........... .. ........................ 20

RECOMMENDATIONS ................. .......................... .. 23

ACKNOWLEDGEMENT .......................... 24

REFERENCES ........................................ ....... 25

APPENDIX A - MODELS, EXPERIMENTAL APPARATUS, AND TECHNIQUE . . .. 27

LIST OF TABLES

Table I - Geometric Properties of NACA 66-210/CCW AirfoilSecticns ............... ........................ .. 33

Table II - Geometric Properties of NACA 64A-212/CCW AirfoilSections ............... ........................ .. 34

~~I1 ~~~ii ;Ii1~.

: J,. : . *.• ,:•,• • •"O-.

LIST OF FIGURES

Fipure 1 - Circulation Control Wing Modes of Operation, PageShort Chord Flap Configuration .... . . . . . . 35

Figure 2 - Extended Trailing Edge Configu:.ation .... ........ ... 36

. Figure 3 - Model Geometry, NACA 66-210/CCW Airfoil . . . . . .. 37Figure 4 - Hodel Geometry, NACA 64A-212/CCW Airfoil ......... ... 38

Figure 5 - 2-D Elliptic CC Airfoil with Spoilers Installed . . 39

Figu::e 6 - 2-D Inserts and NACA 64A-212/CCW Model Installed in8- x 10-Foot South Subsonic runnel ........... . . .. 40 J

Figure 7 - Configuration CCW 243 Installed on Insert Turntableswith Tangential Wall Blowing ..... ............. ... 41

Figure 8 - Aft View Showing Wake Rake, Model, and TangentialWall BloiAng . . . .................... 42

Figure 9 - Lift as a Function of Blowing for the NACA 66-210/CCWAirfoils, q - 30 pef, Rew' 0.64 x 106, h - 0.009" 43

Figure 10 - Effect of Droop on Preventing Leading Edge Separation,NACA 66-210/CCW Airfoil, ac = 8.40, C• • 0.11,Experimental Pressure Distributions ... ......... ... 47

Figure 11 - Effect of Blowing on Experimental PressureDistributions for NACA 66 -210/CCW Airfoil, 6 LE 43.6,tac - -0.86 ............. ... ....... .. . 48

Figure 12 - Effect of Incidence on NACA 66-210/CCW with No LeadingEdge Droop at C• , .088, Experimental PressureDistributions ... . . . . . . . . . . . . . . . . . . 49

Figure 13 - Effect of Incidence on NACA 66-210/CCW Airfoil with2°' Leading Edge Droop at Cu v 0.173, ExperimentalPress'ire Distributions ...... .................. 50

Figure 14 - Lift Curves at Constant Momentum Coefficient for66-210/CCW Airfoils, q - 30 psf, Re 1 0.64 X 106,h = .009". .......... ................... . . . 51

Figure 15 - CEmax and c'STALL as Functions of Momentum CoefficientanA Nose Droop Angle for 66-210/CCW Airfoils ........ 55

Figure 36 -Cj as a Function of Droop and Momentum Coefficient,NACA 66-210/CCW Airfoils ..... ............. .. 56

Figure 17 - Effects of Reynolds Number on Lift Coefficient for NACAk66-210/CCW Airfoil, CCW 800, 6 LE - 00, h - .009" . -. 57

Figure 18 - Drag as a Function of Blowing for the NACA 66-210/CCWAirfoils, q - 30 psf, Re = 0.64 x 106, h - 0.007" 58

Figure 19 - Pitching Moment as a Function of Blowing for the NACA66-210/CCW Airfoils, q - 30 psf, Re 0.64 x 106,h 0.0091 ....................................... 62

iii

PaseK Figure 20 - Lift as a Function of Blowing for NACA 64A-212/CCWUi Airfoils, q - 30 psf, ReX 1.98 x 106, h - 0.027" . . . 66

Figure 21 - Small and Large Chord CCW Airfoil Comparison . . . .. 71

* Figure 22 - Effects of Reduced Trailing Edge Radius - ComparisonBetween Configuration CCW 241 and CCW 244 Experluental

Pressure Distributions ....................... . 72

Figure 23 - Lift Curves at Constant Momentum Coefficient for64A-212/CCW Airfoils, q - 30 psf, Re 1.98 x 106,h = .027' . . . . . . . . . . . . . . . . . . . . . .. 74

Figure 24 - Effective Incidence Corrections for ConfigurationCCW 241 ..... ............ . ............. 79

Figure 25 - Lift Curve Comparison for the 64A-212/CCW Airfoils atCI - 0.16, ag_! 15, h - .027", q - 30 psif .... ...... 80

Figure 26 - Maximum Lift Coefficients within Test Limitations for5 Configurations of a C.C. Wing on a 2-D NACA 64A212L. Airfoil (ag < 15O, h - .027", q - 30 pFf) 81

Figure 27 - Ctmax and Stall Angles for the 64A-212/CCW Airfoils(h - .027", q- 30 psf)... ............ ........ ... 82

Figure 28 - Slot Height and Dynamic Pressure Effects on CI,CCW 241, ag - 9, c - 24" ........ ................ 83

Figure 29 - Slot Height and Pressure Ratio Effects on CCW 241,9° ......................... 84

Figure 30 - Slot Height Effects on CCW 244, ac = ............ ... 85

Figure 31 - Effect of Reynolds Number on Lift, CCW 245(LE - 0%, 6 f - 1800), ac - 00, h - .027" . . . . . . . 86

Figure 32 - Drag as a Function of Blowing for NACA 64A-212/CCWAirfoils, q - 30 psf, Re - 1.98 x 106, h - 0.027" . . . 87

Figure 33 - Drag Coefficient at Cxmax within Test Limits for NACA64A-212/CCW Airfoils, ag < 150, h - 0.027", q - 30 psf 92

Figure 34 - Pitching Moment Variation with Blowing for CCW 241,h - 0.027", q - 30 psf ......... ............... 93

Figure 35 - Pitching Moment as a Function of Lift, Incidence and C1for NACA 64A-212/CCW Airfoils (h - .027", q - 30 paf) . 94

Figure 36 - Pitching Moment Coefficient at Ctmax within Test Limitsfor NACA 64A-212/CCW Airfoils, ag _< 15°, h - 0.027",q - 30 psf ........... ... ...................... ... 99

Figure 37 - Mean Lift and Momentum Coefficients as Functions ofPulsing Rate, CCW 244 ........ ................. ... 100

Figure 38 - Mean Mass Flow Measured by the Venturimeter, CCW 244 . 101

Figure 39 - Corrected Mean Lift and Momentum Coelficients asFunctions of Pulse Rate, CCW 244 ........... 102

iv

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PaeFigure 40 - Effect of Spoiler Deflection on Lift of Elliptical

CC. Airfoil (%a - 0, q - 20 psf, h - .009") . . . . . 103 [Figure 41 - Lift Loss as a Function of Spoiler Deflection . . . . . 104

Figurl! 42 - Reynolds Number and Flow Reattachment Effects onElliptical C.C. Airfoil with Spoiler . . . . . . . . . 105

Figure 43 - Drag Increase due to Spoiler Deflection on EllipticalC.C. Airfoil . .................... 106

Figure 44 - Pitching Moment Variat.on with Spoiler Deflection . . 107

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SYMIOLS

jet exit area. ft 2

Sc airfoil, undeflected chord. ft*

deflected, rotated, or effective chord, ft*

Ca section axial force coefficient

Cd section drag coefficient

Cdrake section drag coefficient from integrated momentumloss on wake rake

f

F C1 section lift coefficient

tt mean section lift coefficient during pulsed blowing

1 clean section lift coefficient of airfoil with spoilerretracted

SCz section maximum lift coefficient

SCLmax aircraft maximum lift coefficient

C1 lift curve slope, per deg~ree

Cm2 5 section quarter-chord pitching moment coefficient

%,n section normal force coefficient

Cppressure coefficient, !.Paq

Cu section momentum coefficient, q-j

C• mean momentum coefficient during pulsed blowing asmeasured by venturimeter

C V2 calculated mean momentum coefficient during pulsedblowing

h jet exit nominal slot height, ft*

ht tunnel height, ft

T measured jet mass efflux, slugs/sec

• unless otherwise noted

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I man jet mass efflox during pulsed blowing, slugs/soc

p local static pressure, pefa

Pd plenum (duct) total pressure, pestdLFd mean plenum total pressure during pulsed blowing, psfa

P maximm plenum total pressure during pulsed blowing,P'max pets uft

Pv line pressure upstream of jjnturimter, psi&

p freestresm static pressure, psfa

q freestream dynamic pressure, pof i

R universal San constant, 1715 ftt/(sec2 R)

r trailing edge radius, ft*

Re Reynolds number based on undeflected chord

S wing or airfoil planform area, ft 2 -

t airfoil thickness, ft*

Td plenum total temperature, °R

Umx maximum velocity in the wall jet velocity profile,ftisec

jV isentropi'% jet velocity, ft/sec

freestream velocity, ft/sec

x distance from undeflected leading edge, ft*

Xslot slot location from undeflected leading edge, ft*

corrected inzidence; ag corrected fot tunt.l boundaryC constraints, deg

aeff effective incidence; ai corrected for induced downwash LIeffects, deg

geometric incidence relative to chord, deg

stall incidence, deg

* unless otherwise noted

vii U

" tI

y ratio of specIfic heats

L 6amean line maximum carber relative to chord line, ft

6f flap deflection * Sie relative to chord, deg

L jet deflection angle relative to chord, dog

leading edge deflection angle relative to chord, deg

68 spoiler deflection angle relative to chord, deg

S£ solid blockage correction factor

X2 two-diuensional airfoil shape factor

P froestream density, slugs/ft 5

density in the jet throat (exit), slugs/ft 3

pulsed blowing frequency, cycles/sec

viii

I '~ SUMMhOARY

Two series of Circulation Control Wing airfoil sections, formed

by the conversion of the sharp trailing edge into a circular bluff surface

with tangential upper surface blowing, were evaluated subsonically to

determine their high lift characteristics as pctential STOL wing sections.

Parameters investigated which had noticeabke effect on the blown Airfoil

Sperformance include leading edge devices (type of device and degree of

defleztton), trailing edge configuration (radius, slot location, deflection,

etc.), Reynolds number, airfoil Incidence, momentum coefficient, slot

height, and nozzle pressure ratio. Maximum lift coefficients rnughly

triple those of the flapped conventionas sections were generated at

incidence slightly less than the conventional stall angles and at blowing

rates obtainable from bleed of state-of-the-art turbojet engines. An

experimental investigation into the lift augmenting effects of pulsed

unsteady blowing was conducted on a smaller radius trailing edge configu-

ration. An additional investigation was conducted to dctermine the effects

of spoilers or similar disturbances ahead of the jet exit. The re3ults

of the above investigations provide a data base for the prediction of the

Rerodynamic characteristics e-f aircraft employing Circulation Control

trailing edges to increase their STOL capability.

INTRODUCTION

Recent investigation into the application to helicopter rotors of

Circulation Control elliptical airfoils employirg tangential blowing

over bluff trailing edges has demonstrated the eimultaneous generation

of hign lift and drag at low blowing rates and low angles of incidence

(Reference ). The concept offers promise of considerable payoff in

STOL performance when applied to the more conventional airfoil sections

of fixed wing aircraft. Based on the well-known Coanda principle, thisCirculation Control Wing (CC4) concept involkes converting the sharp

trailing edge of the high speed cruise airfoil into a rounded surface

to which a jet of air, blown tangentially from the upper surface, adheres.

The Jet sheet is able to turn nearly 180 degrees around the bluff

l I

afterbody, initially acting as a boundary layer concrol, but primarily

cortrolling the airfoil stagnation points and thus the circulation. The

resulting high augiaentation of circulation lift is obtained at values of

blowing or momentum coefficient (C,) well within ihe ranges of mass flow

and pressure which can be bled from state-of-the-art turbojet and turbo-I|fan engines. A problem frequently encountered with blown flap concepts

is the conversion of the blowing momentum into wing thrust, and thus

reduction of drag needed to maintain a low equilibrium approach velocity

at large glide path angles. This is avoided by the Circulation Control

concept, as large jet turning produces viscous mixing with the lower

surface freestream, and jet thrust recovery is significantly reduced.

Application of the CCW to fixed wing carrier-based Naval aircraft

is expected tc9 produce considecable reduction in npproach velocities and/

or landing distances, increased aircraft payload and wing loadings, and

the independence of high speed wing design from compromise resulting from

low speed landing and takeoff requirements. In addition, the generationof lift independent of incidence Implies increased pilot visibility [3

during carrier approach. The existence of an aft upper surface blowing

source provides the additional possibility for increased transonic

maneuverab.Lity. tA characteristic Circulation Control Wing configuration is shown in

Figure 1, where a short chord flap is rotated through 180 degrees to.

expose the cylindrical Coanda surface. Replacement of the conventional

half-span single slotted flap of a 1/5-scale model T-2C aircraft with

a Coanda trailing edge very similar to this configuration resulted in

increases of 96 percent in CLm and 240 percent in the correspondingLax

drag coefficient at C, - 0.156 (References 2 and 3). However, full

rotation of the flap in this configuration results in a loss in chord

of about II percent; additional lift augmentation should result if

this loss can be eliminated. The extended configuration of Figure 2

was thus proposed, where the Coanda surface translates aft to the tr&il-

ing edge as a split flap is deployed to form the lower surface.

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It is the primary objective of the present investigation to provide

section aerodynamic data on the above and additicmal CCW cotifigurations

for use as input to analyses for prediction of three-dimensional perform-

I ance. This report also documents a preliminary investigation conducted

on a small chord airfoil model which was used to cunfirm the effectiveness

of the Coanda trailing edges on relatively thin conventional airfoil

sections. In addition, it was desired to investigate the effects of

unsteady pulsed blowing on the lift augmentation of CCW airfoils, andr.the effects of deflecting a spoiler or similar disturbance upstream of

the blcwing slot. Further goals of this investigation were to identify

a leading edge device capable of preventing laminar separation under

conditions of high circulation, and to examine the effects of Reynolds

number and trailing edge parameter variation on overall airfoil per-

formance.

DESIGN CONSIDERATIONS

Si !Airfoil sections to be modified included those thought to be

characteristic of state-of-thes-art fixed win& aircraft. For the pre-

liminary £nvpstigation on the small chord airfoil, an NACA 66-210

profile was chosen. A larger chord NACA 64A-212 airfoil was constructed

to provide section data to correspond to the T-2C/CCW model bf

Reference 2. Previous 2-D inveszigations of 15-percent thick elliptic

CC airfoils (Reference 4) indicated a serious problem with leading edge

K�searation at high lift, and a primary concern of the 66-210/CCW investi-

/•.•on was the determination of sufficient means to prevent that problem.

A Kruger •eding edge flap and a leading edge droop were alternately

incorporated :ir this purpose. The trailing edge geometries for both

small and large >.ord models were based on those parameters which had

proven effective it IC rotor airfoil evaluations. References 5 and 6sugger that strongly ciýtashed Coanda flow is maintained for a slotheight-to-radius ratio In the range 0.01 < h/r < 0.05, while effective

Coanda jet turning and li•> iugmentation result from a trailing edge

radius-to-chord ratio in the range 0.02 < Y!c' < 0.05. This produces

a slot-height-to-chord ratio of app, :Imately 0.0005 < h/ci < 0.0025,

K3

the lower limit being postulated as the point at which nozzle interial Bboundary layer .thickness approaches the slot height (jet exit thickness)

id seriously degrades the "potential cor'e" characteristics o:- the jet.

Once the trailing edge radius is chosen, the flap rotation point for [the configuration cf Figure 1 is determined by positioning the center of

the radius at the airfoil chordwise location where the sum of radius plus

slot height is equal to local airfoil thickness. This determines the

length of the rotating flap chord and the slot longitudinal location.

For n,ý,qt thinner airfoils, the resulti-. flap chord will be roughly

10 to 20 percent of the original or undeflected chord, c, and the slotlocation ,,ill be approximately .80c to .90c, or aft of 93 percent of Uthe rotatecd chord c' (defilied as the distance from the undeflected

leading edge to the furthest aft point on the roanda surface). It is

essential that this resulting slot location be slightly ahead of the

adverse pressure gradient downstream of the suction peak produced by

attached flow accelerating around the bluff trailing edge. A more

forward slot location can delay section stall due to better boundary

layer control on the upper surface, but a slightly more aft slot produceshigher lift augmentation due to higher jet momentum at the airfoil trail- ii

ing edge. For the extended configuration of Figure 2, the slot lucation

is at the original trailing edge, the effective chord c' is c + r, and

limits on h/r, r/c' and h/ce are the same as above. It should be noted

that all nondimensional data in this report are referenced to the original

undeflected chord c (or original planform area S) regardless of the value 1Jof the effective chord c'; all data are thus directly comparable. Also,

incidence is relative to the chord line through the undeflected leading

and trailing edge points of t'ie conventional airfoil.

MODELS AND EXPERIMENTAL APPARATUS

MODELS

Based on the design considerations of the previous section, two Fairfoil sections were modified into various configurations of the Circu-

lation Control Wing. A small chord (undeflected c - 8.0 inches)

4B

NACA 66-210 section *tan converted into the rotated flap configuration of

Figure 1 and is shawn in Figure 3. KrUger leading edge flaps of 5, 10

and 15 percent undeflected chord were initially employed for leading

edge separation control, but due to their relative ineffectiveness, were

replaced with a leading edge droop, hinged at the lower surface at .15c

and deflected at angles oi 0, 15.3, 29 and 43.6 degrees from the chord.

These configurations are identified as CCW800, CCW815, CCW829, and CCW844

respectively. Additional details of the model are found in Appendix A,iwhile geometric parameters are listed in Table 1.

A large chord (c - 24.0 inches) model NACA 64A-212 was converted to

include the .15c leading edge droop and a number of trailing edge vari-

ations. Figure 4 depicts this model, with additional details found in

Appendix A and geometric patlameters in Table II. Configuration CCW241[ was similar to CCW829, including a 30* droop and 180* rotated flap.

CCW242 was similar except the flap deflection was 900. An extended

trailing edge configuration similar to that of Figure 2 and with 30*

leading edge droop comprised CCW243, while CCW244 was the same as CCW243but with a reduced trailing edge radius having roughly half the value

of rnc'. An investigation of pulsed biowing was also performed with this

Lmodel. Configuration CCW245 was CCW241 with no leading edge droop.A third model was employed to investigatc the effects of upper

surface distortions or deflections of control surfaces such as spoilersSahead of the slot. This model was a 20.6-percent thick cambered ellipse

section characteristic of a rotor blade airfoil, and is shown in Figure 5.A 9.82-percent-chord spoiler with the hinge at 70-1 !rcent chord could be

deflected up to 51 degrees relative to the chord line of the 10.18 inchmodel.

EXPERIMENTAL APPARATUS AND TECHNIQUE H

The small chord 66-210/CCW model was installed and evaluated in the

David W. Taylor Naval Ship Research and Development Center (PTNSRDf"'

15- x 20-inch subsonic tunnel. For this model, as well as all the others,

lift and moment coefficients were obtained from integration of surface

static pressures near the midspan as recorded by a 144- port scannivalve

system. Drag coefficient was obtained from integration of wake momentum

5:

deficit ac measured on a total head wake rAke. Additional details

relating vo model installation; test apparatus and technique; data Lreduction and corrections; and ccntrol of tunnel two-dimensionability

are presci-ted in Appendix A. The 64A-212/CCW airfoil and the cambered

ellipse section were installed in the two-dimensional parallel wall insert

which converted the DTNSRDC 8- x 10-foot subsonic tunnel into a 3- x 8-foot

2-D channel. Figure 6 shows this installation, while Figures 7 and 8 show

configuration CCW243 installed within the inserts, and details of the

wall blowing slots and wake rake. Again, Appendix A presents further

details of installation and technique, and references additional sources

of explanatibn.

A four.:h series of investigations was conducted to determine the

effects of pu.sed (unsteady) trailing edge blowing upon lift augmentation.

Modifications made co the air supply system and data acquisition and

reduction are explained in Appendix A. The model used was the CCW244

configuration (f the 64A-212 airfoil.

RESULTS AND DISCUSSION "

NACA 66-210/C,.W AIRFOIL fAs mentioned previously, the investigation of this smaller chord

airfoil was primarily intended to confirm the effectiveness of Coanda

type trailing edges applied to thin, cambered conventional airfoils for

high lift generation, and to identify satisfactory leading edge devices L. 7to prevent flow separation caused by rapid flow acceleration around rather

sml leading edge radii. The following data shows that, indeed, lack of I.

proper leading edge devices can produce quite severe consequences in lift

augmntatondrag generation and pitching moments.

smaLift Augmentation

Section lift coefficients, as determined from static piessure Liintegration, are presented as functions of momentum coefficient at con-

ýtant corrected incidence (ac) for four different nose droop angles in HFigure 9. Here, the coefficients are corrected for solid blockage and

6

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tunnel boundary interference, but not for any induced effect which mig-/t

Ii result from spanwise non-uniformity due to interaction between trollingedge adverse pressure gradient and tunnel wall boundary layer (see

• lReference 7). The corrected incidence ac has been adjusted to account

for tunnel boundary corrections. At low asd negative incidence, all four

hi-configurations chow a rapid rise in CZ with blowing. This Iritial riseis attributed to flow entrainment and boundary layer control, while the

less steep slope at higher C. is characteristic of stagnation point

movement by the Coanda jet and resulting supercircuLation. However, atlow droop angles, as higher incidence is approached, a noticeable breakV in the curve occurs. The pressure distributions in Figure 10 for ac - 8.4*

and CU • 0.11 very clearly show leading edge separation at 0* droop and

I, reattachment of the flow at higher droop. Furthermore, attached leading

edge flow results in considerably higher trailing edge suction peaks at

a constant CU, indicating greater jet turning, flow entrainment, and liftIV I. augmentation. The inset plot shows an increase of over 1.5 in C1 produced

by applying a 43.6 degree nose droop. As Figures 9 (c) and (d) show, theLi' lift break is delayed by nose droop to much higher values of incidence.

IIInsight into the mechanism of lift augmentation due to blowing is

given by Figure 11. With no blowing, the leading edge stagnation point

is on the upper surface of the 43.60 drooped nose, but as CU is increased,it moves to the lower surface, which approaches full stagnation (C, - 1.0)

along the entire bottom. On the upper surface, a suction spike at the

droop knee is followed by a moderate adverse pressure gradient which

leads, at higher CV, to a long favorable gradient terminating somewhere

aft of 0.97c'. This high upper surface suction and lower surface over-

pressure result in high Cj at quite low Cp. Figures 12 and 13 present

the effect of incidence on airfoil pressure distributions and lft with

and without nose droop. At CU % .088 and 6LE - 0*, Figure 17 ahows

serious leading edge separation at incidence of -0.68 degrees and higher.i This corresponds to the lift breaks in Figure 9 (a). For 29* droop and

CU z .173 (Figure 13), higher incidence produces only slight separation

at the leading edge but the major effect in gradual reduction in the

suction peak at the knee of the droop, and much more pronounced reduction

7

in the trailing edge suction peak. Small changes in the leading edge

flow apparently can quite strongly aifect the boundary layer and down- [stream pressure distributions. As soon as the trailing edge peak begins

to reduce (at incidence above 3.2 degrees for C. - .17) the lift break Uin Figure 9 (c) begins.

Figure 14 presents lift curves at constant Cp, obtained as crossplots

of Figure 9. For all leading edge droop angles, stall occurs at progres-

sively lower incidence with increased blowing; for no droop, astall takes

on negative values. As droop is increased, stall occurs at higher

incidence, and lift augmentation due to blowing is supplemented by lift

due to positive incidence. Figure 15 is a summary plot of this data,showing Ckax and mstall as functions of blowing rate and droop angle.

A comparison is made to the conventional NACA 66-210 with a 600 split flap

(from Reference 8) operatLng at f:Le times the test Reynolds number of

the CCW airfoil. An increase of 250 percent in Ctmax is seen for the

blown airfoil with 6 LE - 43.60 and C. - 0.24; operating at the same LAReynolds number would produce an even larger increase. Figure 14 (d)

shows a similar comparison of the blown and conventional airfoil. Also

shown is the data resulting from the 0.15c Kruger leading edge flap at

Cl, - .142; the loss of effectiveness eue to lack of leading edge flow

control is evident. Figure 16 presents maximum lift coefficient as a

function of CV; in both this figure and Figure 15, it appears th.at the

additional benefits of increasing droop from 29 to 43.6* are minimal.

In fact, at lower blowing, higher droop reduces COmax due to suction on

the lower surface of the leading edge. a% droop of roughly 30 degrees

appears to be more optimum over the entire Cp range, but a more effective

use of droop would be to schedule 6LE as a function of CU and incidence.

However, the pressure distributions of Figure 13 sug:cat Lhat, in spiteof the droop's effectiveness, a device offering stronger control of the [.leading edge boundary layer (such as a slat or tangential leading edge

blowing) might further imnrove the lift augmentation of this type of Uairfoil.

I[38 L

LTnet

Reynolds Number Effects

The majority of th receeding data was run at a nominal freestreamdynamic pressure of 30 paf, yiolding for this small chord model aReynolds number of 0.64 x 106 (based on undeflected chord c and corrected

for model blockage). Figure 17 shows the effects of Reynolds number on

lift for two incidences and two blowing rates for dynamic pressure frbm

10 to 57 psf. At CU - 0, increase in Re produces a slight increase in

C1, but above 0.7 x 106 the difference appears negligible. However, with

blowing, the increase in Ct wtth Re is greater. The magnitude of this

trend with higher CO and Re is not evident from this data. Additional

"evaluation will have to be achieved using the larger chord 64A-212

section data. It should also be mentioned here that maintaining aSL~ constant CO requires higher jet pressure ratios and velocities (Pd/P• and

Vj) at higher Reynolds number. The kinetic energy of the jet (proportional

to V 3 ) and jet turning thus increases with Reynolds number at constant

C1 so that incriments in the Figure 17 curves with blowing may not be due

solely to Re.

In the data of Figure 9, air supply limitations produced an upper

CI limit of roughly 0.12 at q = 30 psf. Additional CU up to 0.20 was

obtained by reducing the dynamic pressure to 20 psf. The flagged data

K ipoints of Figure 9 are obtained at this reduced Reynolds number, but

the curves are faired slightly above these points using the increments

of Figure 17, so that they are indicative of data for Re Z0.64 x 106.

Dra

Drag measurements taken by a rake emlloying 54 total head probes

and 8 static probes were integrated using the methods of Betz and Jones

(Reference 9) and modified (see References 5, 7, and 10) to account forthe additional momentum of the jet so that

* Cdrk CqCd

Cd rake -•rake V4

where 1h is slugs/sec/ft of span in this case.

4_ _9

For most blown airfoil concepts, the jet momentum is converted into

thrust, ane drag is reduced with blowing. As Figure 18 Phows, the same [1is true of the CCW at low blowing rates and moderate incidence. However,

as higher CV is reached, the jet turns further around the bluff trailing

edge and into the oncoming freestream, ptu'ucing a large viscous wake

and increased profile drag. Also, as Figure 18 (a) in particular shows, -

ad soon as the lift break due to leading edge separation occurs, drag

rises very rapidly, even though lift drop-off has not occurred (i.e., I-1$

!o "Cu - stall" has developed). Also to be noted are the high values of

zero blowing drag at low and negative incidence f ,r increased nose drnop,

due priwarily to separation from the underside of the nose. UPitching Moment

As is typical of most blown airfoil schemes, the increased suction

spike far aft on the airfoil generates rather large nose-down pitching

moments about the quarter chord (0.25 undeflected chord). Figure 19

,hows this to be the case with CCW sections as well, with largest values

of nose-down moment occuring for low incidence and high blowing. Higher

incidence of course adds leading edge suction peaks which oppose those at

the trailing 2dge, and m25 becomes less negative when flow remains

attached. A similar result occurs with little or no droop, but is

caused by the separated flow reducing the jet effectiveness of the aft zffsurface and thus the nose-down pitching moment. In general, however,

the conditions of highest jet efficiency are also those of largest nose-

down moment.

NACA 64A-212/CCW AIRFOIL flBased on the initial confirmation by the 66-210/CCW section of high

lift and drag generation at. low incidence and blowing, and on the identi-

fication of roughly 300 droop as an effective leading edge device, the

64A-212/CCW section was tested to investigate: additional trailing edge

configurations; effects of higher Reynolds number and slot height

variations; and the aerodynamic properties of the airfoil section employed

in the T-2C/CCW investigations of Reference 2. The configurations have [3

10 l

~" ,-4~- IIlk

[SI been identified in the pievious section on models ard in Figure 4, and.• the geometric, parameter@ are listed in Table 11.

Lift Aumeantation

Figure 20 presents lift as a function of Cu at constant corrected

incidence for five configurations, all run at a dynamic pressure of 30 pef

and slot height of 0.027 inches. The lift break due to leading edge

problems at higher i-tcldences on the 66-210/CCW is still observed,although for similar configuration (CCW829 and CCW241) it now appears to

be oc.urring at much higher incidence land greater CE for the larger chord

model. Figure 21 shows a clearer comparison of these airfoils. At zero

incidence, the relatively small difference in C1 is probably due to

Reynolds number, while at 9" geometric incidence, the much larger differ-

enc-s are probably due to more severe separation on the smaller leading

edge radius of the 66-210 airfoil. Leading edge separation control on

CCW241 allows this airfoil to generate Cj - 5.75 at CU - 0.19 and mg - 9*

with a chord that is shortened 11.4 percent from the undeflected value.

"The pressure distributions for the 64A-212 airfoils show trends very

similar to Figures 11, 12, and 13 for the smalleý, airfoil. In Figure 20 (e),

CCW245 (no leading edge droop) displays separation problems similar to CCWSOO

in Figure 9 (a). Increased Reynolds number and nose radius appear to

* produce little improvement for the undrooped airfoils.

Configurations 242, 243, and 244 display the effects of trailing

K edge geometry variation, Figure 20 (b), (c), and (d). The 90' flap, an

intermediate rotation angle for the shortened chord configuration ofFigure 1, generates high C, due primarily to the large effective cainber

and the fact that no chord loss exists in this partially rotated configu-

ration. However, at higher C,,, operation at incidence above 0 degrees

produces stall (i.e., decreased Cj with increased a). The reduced radius

trailing edge configui tion (CCW 244, r/c' - .021) shows rather interesting

trends when compared to the larger radius (r/c' - .041) of CCW241 for

incidence of 6 degrees and less. Figure 22 (a) indicates that for no

blowing and mg - 30, the aft lower surface camber produced by the rotated

flap produces add tional Cj for the larger radius trailing edge. However

11.i I

at higher C., the reduced radius trailing edge oustains pressure coef-'

ficients roughly twice that of CCW241 at the same blowing; this increased

let turning and entrainment produces significant additional C1. At

higher incidence, Figure 22 (b) shows the larger radius produces greater UCZ in all cases. The aft camber predominates at low blowing, and the

slightly farther forward slot (in terms of actual deflected chord) of L4CCW241 provides better upper surface boundary layer control at high

incidence and higher C., even though the smaller radius still provides -

better jet turning. The extended trailing edge configuration (CCW243 H

of Figure 4) shows significant lift increase over CCW241 by regaining Ii

the chord lost during the airfoil's flap rotation. A Cu limit of

roughly 0.08 was imposed by the model's long unsupported upper plenum

surface between the adjustment screws; this expanded under pressure and -

consequently closed the blowing slot. However, based on similarity to

CCW241, it is felt that the extrapolation to higher Cu in Figure 20 (c)

is valid.

The more conventional lift curves (C1 vs a at constant CU) are

plotted in Figure 23, where the conventional 64A-212 data from Refereuce 8

is compared in Figure 23 (a). All configurations except CCW245 produce

increases of at least 215 percent in C1ax at CU - .16 over the conven- Utional split flap airfoil at Re - 6 x 106. Configurations 241, 243, and

244 all posses3 the conventional lift curve shapes with parallel slopes

and positive stall an8les, decreasing somew¢hat with increased Cu

(analogous to in..rea-ed conventional fiap deflection). However, CCW242 Uand 245 possess low or negative stall angles ý higher C., both apparently

expectencilig leading edge problemo. The 90* flap data shows very little

parallelism in lift curve slope. ThE operation of both airfoils seem to

suggast problems In areas )f posLtive a-operation (-uch as gusts or Liconventional land~ng /lerar) and in systems to control blowing u.id

incidence scniedW--% in these regions of operation.

In Figure 23, additional dashed curves are provided which correspond

to t\, actual effective incidence, aeff' at which the airfoil nperates.

These values are obtained from curves similar to Figure 24 for each air-

foil. The effective incidence is ac modified as described in Appendix A

12 i

to account for the induced downwash field due to vorticity shed at the

* I tunnel wall intersection with the modal, and is a function of lift

coefficient (i.e., trailing edge suction puak). As can be seen, the

corrections are significant. At xero blowing, aeff =c since no trail-Lcing edge peak is present. Using the effective incidence, Figures 25 and

26 sumurize the data nf Figure 23 in terms of Ct obtzined at a higher

.ivalue of CU over the aeff range, and C#,max as a function of Cu. Figure 27

plots Cimax versus effective stall angle for the five configurations at

Swvarious values of Cu.

Slot Height Variation

The momentum coefficient is a function of the parameters Pd/P.' Td,

Ap, S and q and for a two-dimensional airfoil may be defined as

iv~ J pjhVJ2 0 h VJ 2

.j.0V 2c c

Figure 28 shows experimental vas .ation in the momentum coefficient for

configuration 241 at 9* geometric incidence. Although theoretical Cu

is not a function of incidence (see Appendix A), when mass flov is

measured experimentally the significant pressure ratio is total pressuredivided by local static pressure at the slot exit, which will of course

very with q, airfoil configuration, lift coefficient and incidence.

Figures 29 and 30 show the effects of slot height variations on

I lift for CCW241 and CCW244. Both figures show that for a constant value

of CU, greater lift results from a smaller slot. This is of course due

to the fact that the reduced slot area is compensated for by increasedjet velocity to maintain a constant CU. Increased Vj immplies increased

kinetic energy, jet turning and lift augmentation. However, these

figure. show that there are crossovers where reduced slot height corre-

sponds to reduced C1 at constant CU. It will be noted that these values

of h/c' approach the lower livtt of .0005 suggested in the Design Con-siderations section, where internal nozzle boundary layer thickness

13

('b

begins to equal the slot height. Figure 29 also indicates lines of -

constant pressure ratio for tha two larger slot heights; these lines irepresent constant jet velocity. In this case, qn increase in slot

height produces ?nstreased lift due to the resulting higher mass flow iiand Cu. The finae decision conce.ning use of reduced or increased

slot height will capend on the method of air wupply: a fixed pressure Uratio system would suggest increase in slot height, wherea& a constant

Cp system might suggest reduced slot height for improved performance

(although the required power, proportional to Vj 3 , would be an inportant

consideration).

In contrast to the detrimental effects of too small a slot height,

Figure 30 also confirms the problem of partial jet 4etachment with large

slots (see Reference 11 for additional details). Under certain conditions

of high pressure ratio, large slot height, and reduced trailing edge

radius, the Coanda mechanism begins to break down, and the jet turning is

considerably reduced, if not totally eliminatedby jet detachment from

the surface. The higher Cu data points for h - .054" show significant

scatter and indications of jet ineffectiveness; pul3ating turbulence

from the vicinity of the model confirmed this. The parameter hic' is

at the upper limit suggested in the Design Considerations. Thus, both Iupper and lower limits on slot height are confirmed, but the associated

values of C. which should not be exceeded at those parameters have not hbeen established for all configurations and incidences.

Reynolds Number Effects

Use of the larger chord model allowed Reynolds number variation ur

to 2.4 x l05, corresponding to q - 50 psf. The effects on lift are given

in Figure 31 for CCW245 at ag - 00. Again, low C. produces relatively

little Ck variation with Re, but at higher C. a rather noticeable increase

is present. The effects of Re alone cannot be distinguished from the

effects of increasing the pressure ratio (jet velocity) in order to main- Htain constant Cp at increased q. It is suggested that lines of constant

14

!I

C1mj, C IjPd/P. or C, 0 4 parate the Reyr.dbids number and jet

"velocity effects.

Results of variations In section drag coefficient with blowing in

Figure 32 are very similar to those of Figure 11 for the small chord

airfoil: (1) nose droop produces high Cd at no blowing and low incidence,

(2) application of blowing brings on rapid drag reductions at all

incidences, and (3) viscoas mixing on the aft lower surface or leading

edge separation result in rapid drag increases with increased C.. A

.number of the data points found in Figure 20 have had to be eliminaczed

from the drag plots due to movement of the momentum deficit region of

the wake off the lower limits of the rake. In addition, data scatter at

higher CV is attributable to more sparsely located total head probes

near this lower limit of 'he rake. The absence of drag eata for CCW242

at higher CV is due to large regions of turbulence behind the 15% chord

flap perpendicular to the flow. (This large turbulent wake could produce

serious problems when applied to a finite wing aircraft if the wake

should fall on the horizontal tail surface.) With the 7bove limitations

L. in mind, a comparison of the drag data for the five configurations at

C~max points of Figure 26 is made in Figure 3.3. With the exception of

the 90* flap case, all configurations show a rapid decrease in Cd with

CU, followed by either a drag rise or a change in slope which indicates

the beginning of a drag rise. In contrast, CCW242 shows constantly

decreasing drag with no apparent change in sJ., indicative of a drag

rise; this is due to the fact that Cimax for this airfoil always occursI at relatively low incidence, and to the fact that the jet can turn no

further than 90" due to the existence of the flap surface. A character-

istic Jet flap mode with a higher degree of thrust recovery is evident.

Pitching Moment

The variation of quarter-chord (25 percent undeflected chord)

pitching moment with CV at constant ac is presented for CCW241 in '1Figure 34. As is charac:;.ristic of almost all high lift schemes, the

Lb {15

suction peak produced by the trailing edge device (be it blowing or a

flap) generated additional nose-down pitching moment. There appears to

be little variation with incidence (except above 12"), Indicating that at

constant C. the trailing edge suction peak is dominant over the variation

in leading edge peak due to incidence. Figure 35 presents C1 vs C

along lines of constant incidence and CM for all five configurations. Thet11

larger nose-down moments associated with the 90" flap case are due to the

near vertical component of jet thrust acting parallel to the flap surface Hwith the high thrust recovery characteristic of jet flaps. This thrust

contribution to moment is considerably reduced for all the bluff trailing

edge configurations due to jet mixing. However, the exterded trailing .

edge configuration (CCW243) does exhibit somewhat higher moments than the

other bluff versions in that its jet suction peak is located near tCe IIundeflectod airfoil trailing edge, roughly .15c further aft. Figre 36 1points out these relative comparisons by plotting the pitching moment Lcorresponding to Ctmax at various Co. In addition to the above explanation,

the greater nose-doon moments of the 90* flap also occur because Ckmax

for this airfoil is always at low incidence, where the more favorable aft

pressure gradients serve to create even higher jet suction peaks over the "4

Coanda surface, and where th- counter-acting leading edge suction peaks

are less.

PULSED BLOWING INVESTIGATION

Investigations conducted with pulsed unsteady blowing over a blown Kflap (Reference 12) and on a cambered CC ellipse (Reference 13) indicated

that definite additional lift augmentation above the steady-state value

coald be obtained by this technique. For equal values of effective H *i

momentum coefficient (time-averaged C1 in the unsteady case), the pulsed

blowing prodnced higher trailing edge suc:ion peaks and lift augmentation LIbecause of the instantaneous high... values of duct pressure and jet velocity

which produced greater flow entrainment and jet turning. This of course ii!iallows a desired lift coefficient to be produced at reduced mass flow rates. ' VThe results of Reference 12 indicate this reduction in required it to be as U

much as 50 percent. Both investigations identified optictum pulsing fre-

quencies, above or below which lift generation was less.

16

S. . .. .. . . . . I

A similar investigation was conducted on configuration CCW244,

.i " which has the reduced trailing edge radius (r/c' - .021), to determine

if similar reductions in required mass flow could be realized for a

K :configuration of the CCW. Appendix A describes the test setup and data

reduction techniques employed; the majer difference from the previous

tests was the insertion of a rotating butterfly valve between the venturt-meter used to measure system mass flow and the plenum of the model.

Figure 37 plots averaged lift and momentum coefficients av functions of

KI pulsing rate. The lift Is averaged automatically by the damping of the

static pressure tubing, whi:e mass flow is averaged by recording mean

.I readings of the venturimete.' pressures. Duct pressure and jet velocity

are obtained by identifying mean values of Pd from oscillograph traces of

K the model plenum jressure transducer output. There is no apparent trend

in lift augmentation as a function of pulsing rate from 0 to 69 cycles

per second in Figure 37, with the exception that all unsteady data appears

to be less effective than the steady case (w - 0 cps). It was suggested

ni by the lack of any smooth trend with variation of w that some error was

being introduced by the venturimeter. Figure 38 plo~s mean mass flow

recorded from this flow meter as a function of line pressure upstream of

.�the pulsing valve. As data from Reference 12 confirms, a given mass flow

at the slot should always require higher upstream line pressure than the

Ksteady cate, for which the cyclic valve is producing no pressure loss.

However, as Figure 38 shows, mean mass flow for the 20 and 40 cps cases

is produced at lower line pressure than the steady case. It was concluded

that the venturimeter was being adversely affected by the pressure

fluctuations feeding oack upstream from the butterfly valve, and its

measurements were unreliable.

As an alternative, mean duct pressure Pd read from the oacillograph

traces was used with a previous calibration to determine slot height

deflection under pressure, and c6. resulting slot areas and pressureratios were substituted into th.e isen'ropic equations in Appendix A to

yield mean maas flow, jet velocity and CW. The resulLs in Figure 39

inoicate that for the cases investigated, pulsed blowing has littlediscernable lift augmenting effect. This may be due to two aspects of

17

the test setup. The pulsing device of Reference 12 produced a square

wave variation of duct pressure in which the minimum was almost zero--

this produced a maximum value almost twice the mean, and thus jet turning

and augmentation that were considerably higher than for the mean CW. The

pulsating valve used for the present investigation produced a si, isoidalpressure variation in which the maximum pressure was never more than U15 percent higher than the mean or comparable steady state pressure. Thus

th. augmentation was minimal. The comparison is sketched below.

At dE

max

i, d Pdidma

'• . . .. .. .. . ' " • • .

time time

Reference 12 Present Investigation

Secondly, the trailing edge configuration was probably not ideal for this LIinvestigation. The small radius (r - .4375") and larger slot produced

h/r - 0.0617, which under conditions of highcr jet velocity could produce Udetrimental effects. Furthermore, as Reference 12 indicates, the vorticit-

produced by jet pulsing which is responsible for the additional augments-

tion begins to decay when the time intervals between the pulsing become

greater than approxiv Ltely 50 r/Ureax, where Umax is the maximum value in

the wall jet velocity profile. Thus small radii are detrimental to

unsteady augmentation under higher jet velocities or greater p:essure

ratios. Therefore,concidering this along with the ineffective pulsing

valve used, the present investigation did not verify the positive effects

of pulsed blowing determined in References 12 and 13, but did suggest

more appropriate model and valve conifigurations for any further

experiments.

18

7-__

SPOILER INVESTIGATION

U The application of CCW airfoil sections promises to lower the

ui approach and take-off speeds of fixed wing aircraft to the point where

U conventional aerodnamic control devices may beccme ineffective.

Reference 2 data shows that ailerons on a partially blown wing can be

LI exposed to severe spanwise flow or outboard Reparatton, and in addition

can generate adverse yawing characteristics. It was surmised that

spoilers, located upstream of the blown trailing edge, would be a far

more effective roll control and would not produce the adverse yaw

problem. The possibility of a spoiler Ibeing too effective when deflected

into the boundary layer just upstream of the tungential jet led to a

preliminary two-dimensional investi.gation. The model, a 20.6 percent

thick ellipse with 3.6 percent circular arc camber and r/c - 0.045 had

previously been installed in the 3- x 8-foot inserts, and was felt to be

H characteristic of the trailing edge region of a CCW profile even though

it was a rotor blade section. Figure 5 depicts the model and the two

spoiler configurations employed. Both were -Liads from angle iron to pro-

vide rigidity over the 36-inch span, and were located so as to pivct about

a point on the upper surface at 0.70c. The spoiler length cs was 0.0982c,

"and deflection relative to the airfoil chord was from -9* (flush to the

surface) to +51@. The slotted version of the full spoiler was constructed

.Jby removing half the spoiler length nearest to the airfoil, so that part

of the airfoil boundary layer flow could penetrate through to the jet slot.

H IAll tests were run at zero airfcil incidence at q - 20 psf (Re - 0.66 x 106).

Wall blowing to prevent three-dimensionality was not employed. Data are

corrected for tunnel blockage and boundary constraints, but not for induced

incidences.

Figure 40 presents lift as a function of blowing for various spoiler

deflections and types, and for the clean airfoil (6, a -9°). The effect

on C1 of deflection of the full spoiler is quite severe; supercirculation

V is heavily restricted and the airfoil behaves quite like a jet flap, with

lift almost linearly proportional to blowing. Even with the small spoiler

K i deflection of zero degrees (a perpendicular aft-facing step of only

3 percent chord), lift at C1- 0.16 is reduced to 28 percent of the clean

LI

U - 19J

airfoil value. The slotted spoiler shows some improvement: lift returns

to the more conventional form where Ct - f(4nv and supercirculation is

partly restored. Figure 41 plots lift loss (CZ CI - CzcIean) as a Iifunction of spoiler deflec•ion and C1. The full spoiler possesses the

undesirable trait that lift loss is not linearly proportional to spoiler

deflection; the majority of reduction occurs bet-#een -9o and 0. The

slotted spoiler is more linear and sbould have more desirable control

characteristics. UFigure 42 shows that at an increased blowing rate, the separated

flow behind the full spoiler at 6. 0* can be made to reattach due to

increased flow entrainment at higher Cu. The effect of Reynolds number

is seen to be minimal for this case, at least below the reattachment 1point, but there appears to be some Reynolds number effect on the clean

airfoil.p ~The increase in drag due to spoiler deflection is evident in LFigure 43; a smaller increase in drag occurs when using the slotted

spoiler. Similar trends in pitching moment are seen in Figure 44; nose- Iidown values are greatly reduced by spoiler deflection. Both of these

--ends may provide effective control when applied to fixed wing operation, Halthough the non-linearity with 6. may present a control system design

problem. Furthermore, it appears that the slotted spoiler at relatively

low deflection should provide a mnre satisfactory control system for the

C04. Reduced lateral span of the spoiler might reduce the roll sensitivity

to deflection, but this parameter and deflection schedules will have to Libe determined relative to a specific three-dimensional configuration.

';SIONS

A esries of subsonic two-dimensional investigations conducted on

a number of configurations of Circulation Control Wing airfoil profiles

has confirmed the ability of thin sections employing the concept to

generate high lift at relatively low values of momentum coefficient and

incidence. A data base has been established for use in predicting the

aerodynamic characteristics of airt:aft with CC trailing edges, and the

effects of variation of a number of aerodynamic and geometric parameters Li

20

ULon airfoil performance have been identified. The experimentsl data

obtained from the four investigations yielded the following conclusions:

NACA 66-210/CCW Airfoil (tested over a range of a, CV, 6 LE, and Re)

LI A 0.15c leading edge droop of 43.6* and Cu - 0.24 produced a

section rc - 5.5 and satisfactory airfoil operation at higher

Li positive incidence (uncharacteristic of CC rotor •.ade sections).

Reduced droop angle resulted in leading edge separation, reduced

Ui jet turning, strong breaks in the CQ vs CU curves, rapid draS

rise at onset of separation, and failure of the ction to operate

[L effectlvely at positive incidence. Similar resuits were observed

for KrUger leading edge flaps.

e Effects of Reynolds number on lift were minimal with no blowing,but produced positive lift increaents with increased Re at

constant Cu.

0 o For attached leading edge flow, increased CU initially reduced

drag but eventually caused a rapid drag increase as lower surface

i visrcous mixing occurred. Separated flow at the leading edge

caused immediate drag increase.

* Pitching moment rapidly became more negative with blowing, and

took on the largest negative values at reduced incidence and

[ i greatest jet effectivene3s.NACA 64A-212/CCI Airfoil (tested over a range of a, Cu, 6LE' h, Re and

trailing edge configuration)

F The 2-D section representing the reduced chord T-2C/CCW airfoil

Hyielded CZ - 5.95 at - 0.20 and c- +9; this was extrapolateato C1 - 6.30 for the extended trailing edge configuration at the

U same blowing and slightly lower incidence.L The greatest net lift was generated by the 900 flap configuration

due to its full flap chord and high effective camber, but it

i displayed a number of disadvantages subtracting from its suita-

bility. These include very large nose-down pitching moment,

i large turbulent wake, low or negative stall angles and negative

lift curve slope at most positive incidences.L..LI 21

>_____

oA reduction of 502 in the trailing edge radius of the shortened

chord airfoil produced comparable or better lift performance due

to very high jet effectiveness. Its performance was slightly

reduced compared to the larger radius at higher incidence due to

its slightly further aft slot location in terms of effective

chord. LDrag, pitching moment, and Reynolds number trends were similar to

* ~those of the 66-210/CCW airfoil, with the exception of the 900 Eflap discussed above.

oSlot height increase can 'have differing effects on lift augmen-

tation, depending on the mode of operation of the blowing system.

Increased h yielded reduced CX~ for constant values of C.~ due to

reduced jet kinetic energy, but increased CX resulted if the Ipressure ratio was retained constant and h increased. Suggested

upper and lower limits on slot height were confirmed, beyond which

adverse effects on lift augmentation occurred.

Pulsed Blowing Investigation (variation in Cu and w)

AResults showed that rio benefits were gained from unsteady blowing

as performed. This was due to the poor response pattern of the

valve and a small trailing edge radius on the airfoil.

o A valve capable of a square wave variation of Pd with time

(resulting in much higher Pdmax for a given mean duct pressure)

and a larger trailing edge radius appear necessary to realize

any lift augmentation due to pulsed blowing.

Spoiler Investiga' -. ons (variation in CV, 69 and spoiler type) .

oThe full spoiler showed very large reductions in C1, Cd and nose-

down pitching moment as compared to the clean airfoil. Allfl

reductions appeared to be non-linear with spoiler deflection.

* .. . aThe slotted spoiler shmwed reductions in lift, drag, and pitching

moment not quite as large as those for the full spoiler, and

the variations appeared to be much more linear with deflection.

{ -~---....22 Ln' .7.

RECOMIEWDTIONS

LI The following recomendations are suggested by the present

investigations and by the listed conclusions:

Li *Whereas leading edge droop appeared to control leading edge

separation, it was detrimental at lower incidence and appeared

to still allow some undesirable boundary layer development it

high incidence and lift. A more effective leading edge device,

such as a slat or leading edge tangential blowing, should ba

substituted for droop in a future experiment.

The significant performance variations with trailing edge

configuration suggest that other geometries might prove more

K optimum. The deviation from circular Into cycloidal or other

geometrically simo.lar shapes is suggested. A parametric study

of trailing edge variation needs to be undertaken, and thedevelopment of an analytic method able to predict the viscous

characteristics of this type of airfoil is strongly suggested.

To further investigate the possible advantages of pulsed blowing,

an additional 2-D investigation is recomended with a larger

U trailing edge radius and an alternate pulsing mechanism which

produces larger Pdmax relative to the mean duct pressure.

Additional spoiler investigations are suggested where small

deflections in the range of -9 to 0 degrees for the full spoiler,

Li and -9 to 25.5 degrees for the slotted spoiler are evaluated

for possible linearity in those ranges. Also, a spoiler on a

i 3-D configuration should be investigated in which variations in

spanwise length and lateral location of the spoiler are

I conducted.

4ULi

l23I-

ACK(NOWLEDGMENET

The author would like to express his appreciation to the following

individuals for their assistance in conducting the wind tunnel investi-

gations reported in this study:

Mr. Phil Dodd for assistance in tunnel operation and data

recording during the 66-210/CCW investigations in February

through May, 1972, and Mr. Jonah Ottendoser for design,

installation, and testing of the KrUger leading edge configu-

ration of this model in the spring of 1971. [3

Messrs. Jonah Ottensoser, Jerome Winkler, and Dulany Davidson 11for assistance in model installation, tunnel operation and

data collection during the 64A-212/CCW investigation in May Uand June of 1974.

Mr. Fred Krause for similar assist.ance during the pulsed

blowing investigation in August and September 1974.

Mr. Rod Hemmerly for similar assistance in the spoiler tests

in March 1975.

V

24! I iU

il • 2lii

REFERENCES

1. Stone,.M. B. and R. J. Englar, "Circulation Control A

Li Bibliography of NSRDC Research", July 1973, Naval Ship Research and

Development Center Report 4108.

2. Englar, Robert J., "Subsonic Wind Tunnel Investigation of

Ki the High Lift Capability of a Circulation Control Wing on a 1/5-Scale

T-2C Aircraft Model", May 1973, Naval Ship Research and Development

Center Report ASED 299.

3. Englar, Robert J., "Investigation into and Application of the

Li High Velocity Circulation Control Wall Jet for High Lift and Drag

Generation on STOL Aircraft", AIAA Paper No. 74-502 presented at AIAA

V 7th Plasma and Fluid Dynamics Conference, Palo Alto, Calif., June 1974.

H 4. Englar, R. J., "Two-Dimensional Subsonic Wind Tunnel Tests of

Two 15-Percent Thick Circulation Control Airfoils", Aug 1971, Naval Ship

Research and Development Center Report ASED 211. (AD 900 210L)

S5. Englar, R. J., "Two-Dimensional Subsonic Wind Tunnel Tests of

a Cambered 30-Percent Thick Circulation Control Airfoil," May 1972, Naval

Ship Research and Development Center Report ASED 201. (AD 913-411L)

6. Englar, R. J. and R. M. Williams, "Design of a Circulation

Control Stern Plane for Submarine Applications", Mar 1971, Naval Ship

Research and Development Center Report ASED 200. (AD 901-198)

7. Englar, R. J. and R. M. Williams, "Test Techniques for High

Lift Two-Dimensional Airfoils with Boundary Layer and Circulation Control

for Application to Rotary Wing Aircraft", Canadian Aeronautics and Space

Journal, Vol. 19, No. 3, pp 93-108, Mar 1973.

8. Abbott, Ira H. and Albert E. Von Doenhoff, "Theory of Wing

t { Sections", N.Y., Dover Publications, Inc., 1959.

I25

9. Schlichting, Hermann, "Boundary Layer Theory", 6th ad., N.Y.,

McGraw-Hill, 1968.

10. Kind, R. J., "A Proposed Method of Circulation Control", Cambridge,

Eng., June 1967 (Cabibridge Univ. Ph.D. Thesis).

11. Englar, R. J., "Experimental Investigation of the High Velocity LiCcanda Wall Jet Applied to Bluff Trailing Edge Circulation Control Airfoils",

June 1973, Naval Ship Research and Development Center Report ASED 308.

(M. S. Thesis, Dept. of Aerospace Engineering, University of Marylaad, 1973).S~LI

12. Williams, J. R., et.al.,, "Analysis of a Pulsing Wall Jet", Oct 1972,

Colombus Aircraft Division, North American Rockwell RepoA-t NR72H-325. [j13. Walters, R. E., et. al., "Circulation Control by Steady and

Pulsed Blowing for a Cambered Elliptical Airfoil," July 1972, West Virginia

University Department of Aerospace Engineering TR-32.

14. Englar, R. J. and J. Ottensoser, "Calibration of Some Subsonic

Wind Tunnel Inserts for Two-Limensional Airfoil Experiments", September,

1972, Naval Ship Research and Development Center Report ASED 275. &on

15. Pope, Alan, "Wind-Tunnel Testing", 2nd ed. N. Y., John Wiley Sons,

Inc., 1964.

26

* -.

. . . . . . . . . . . . . . . . . . . . .

APPENDIX A A

Li MODELS, EXPERIMENTAL APPAPATUS, AND TECHNIQUE

MODELS

Both CCW modifications of conventional airfoil sections were based

on the geometric parameters of Tables I and 1I and are shown in Figures 3

and 4. Coordinates of the airfoils ahead of the slot are identical to

the original airfoils with the exception of deflected leading edge devices.

Leading edge droop for both airfoils involved rotation of the nose about

a lower surface hinge point at 15 percent of the original chord, and

filling the resulting gap with a circular arc with center at the hinge

point. Slot height was adjustable by means of compression and Jack screws;

the slot was contoured to be the minimum area of a smoothly converging

nozzle. Nominal slot height is the value measured with feeler gage stock• with zero plenum pressure. Slot expansion with pressure was measured for

each configuration and found to be linear for the 66-210/CCW airfoil, but

of no specific trend for the 64A-212/CCW airfoil since the plenum geometry

was not always identical for all configurations. On the 6-210/CrCW Rection,

the trailing edge was fixed in the fully rotated position of Figure 1 and

only the slot height and droop were variable. For the 64A-212/CCW configu-

ration, the trailing edge confijurations were interchangaable as shown in

Figure 4, and droop and slot height were variable. The CC rotor airfoil

used in the spoiler test (Figure 5) was built by Lockheed-California

Company under contract for application to & research rotary wing aircraft;

for the present tests, only the deflection of the added spoilers was

variable. All models had a chordwise distribution of static prossure taps

at or near the midspan, and one to three rows of spanwise static taps to

measure two-dimensionality from wall to wall. Plenum total pressure and

temperature were recorded by a total pressulre probe and thermocouple

located in the plenum near midspan.

L27

_16i N V .4K e,4A:-

TEST APPARATUS AWT' TECHNIQUE

The small chord model 66-210/CCW airfoil was installed in the LDTNSRDC 15- x 20-inch subsonic model tunnel, spanning the short dimension

of the tat section and attached to circular wall mount. which rotated to

vary incidence. All pressure data from the airfoil and wake rake wam

recorded automatically by a 144-port scannivalve system, punched on paper

tape, and reduced by XDS-930 di$ital computer. Mass flow into the modei

was measured by orifice plcte or venturimeter in the supply line. Test

technique was identical to that employed in previous tests of elliptic

rotor blade airfoils (Reference 1) and presented in detail in Reference 7.

One exception was that the strong three-dimensional induced effects pro- Uduced by interference between wall boundary layer and jet adverse pressure Ligradient were controlled by tangential blowing from the tunnel vertical

sidewalls at stations ahead of the nose and on the uopper surface ahead of

the Jet 3uction peak. Blowing rates from these jets were set to those

values which best restored two-dimensionality to the spanwise pressure

taps as displayed on manometer boards.

Both the 64A-212/CCW and the elliptic airfoil were installed on

rotatable endpi.ates in the walls of the 3- x 8-foot 2-D inserts installed

in the DTNSRDC 8- x 10-foot subsonic South Tunnel. Figure 6, 7, and 8

show this installation, and Reference 14 describes in detail the develop-

ment, calibration and operation of this facility. Data recording, mass

flow measurement, and test technique were similar to the small tunnel

investigations, with the exception that 3nly one set of wall blowing

slots was used to control two-dimensionality due to limited facility

air supply. It was found later that these were not totally effective,

and it was necessary to correct the data for induced effects, as discussed

below. For the spoiler tests, no wall blowing was employed at all, since

only data trends were of interest there.

DATA REDUCTION AND CORRECTiONS 1)As mentioned previously, all force and moment data was from inte-

gration of model static pressures or deficits in wake momentum distribu-

tions. All coefficients have been nondimensionalized by the original

I2828 I

S ' .. . ... ... .. . . . . ... , , • . . . " ' ' " . . .. . . . . . . .. . . . . . .I•

4Lhord which connects undeflected leading and trailing edges (8 inches

for 66-210, 24 inches for 6W-212, and lOb Inches for the cambered Vellipse). Incidence is referenced to this chcrd, and Reynolds numberis based on it as well. This allows direct comparison of all airfoil

results. It should be noted however, that the pressure distrihutionsfor the 66-210/CCW airfoil are plotted in terams of x/c' to stretch out

the longitudinal coordinates. Normal and axial force coefficients from

pressure integration (subscript "p") were modified by the jet thrust

component and then combined to yield the lift coefficient:

Cr W Cnp CP sin 6 where C p -p dcp fp

a - C coo 6 where C pP p'

CQ. n CO asc Ca sin ac

' Vwhere 6j is the angle between the chord and line of jet thrust, and Mc is

the corrected incidence. The drag integration employs the methods of

Betz and Jones (Reference 9) to determine the drag coefficient Cdrske,whicn must then be modified (References 5, 7 and r0) to account for the

additional jet momentum in the control volume:

;IV.. VMK C�Cd -C

rake qc drake

The pitching moment coefficient Cm2 5 is determined from integration of thenormal and axial components of static pressure and the jet thrust reaction

about the quarter-chord point. The freestream dynamic pressure is

corrected for solid blockage as follows:

29

S• -[1, , , 29- : ,

q qu (1+2c)

where c - 0.822 A (t2ht) 2 2

and X is an airfoil shape factor from Reference 15. HThe above coefficients are referenced to this corrected value of dynamic Lpressure, and corrections for tunnel wall constraint (induced camber) are

then applied (where subscript "u" is the uncorrected value as determined

above) :

C = •. -ICQ~m E Su 92 \tt u.

Temomentum coefficient was c~alculated as

m• =n (if t•t is per unit sipan)M2 qS qc

where the total mass flux (&) was recorded experinentally from the install-ed flowmeter, and the jet velocity was calculated assuming an isentropic

expansion from plenum total conditions to freestream static condition:

VJ2w2RTd/ y i I P

30

For the Dulsed blowing cases where it was necessary to calculate the

isentropic mass flows, these were determined as follows:

choked flow: 1; Aj Pd

Fi (,r)d

•' For the 64A-"i2iCCW airfoils, one additional set of data ccrrections

,•, was applied. In this case, the tangential tunnel wall blowing did r~ot

4 prove totally effectivh in eliminating the three-dimensionality andi [ induced effects discussed in the text, and it was evident that the airfoil

was operating et an effective incidence less than •c' As discussed in

+ !' iReferences 5, 7, and 10, this effective incidence could be determined bySL_ matching potential flow presmure distributions with the experimental

distrabutions It the same effective Cg over a range of incidence, andpdeterotining that incidence at which the theoretical and experimental laad-

inug edge stagnation points coincided. This would bevh thtrue effectiveIw incidence of the experimental data. With increasing C., this incidence

becomes considerably less than efct Incidence correction curves similar

to Figure 24 were calculated iiw this manner for each ccnfiguration of

the 64A-212/CCW airfoil, and the lift curves of Figure 23 were corrected

I~i accordingly.

PULSED BLOW~IING APPARATUS

deteFor this investigation, a housing containing a rotat'ing flat disc

driven by a variable frequency motor was installed in the air supply line

i droughly 8 feet downstream of the venturhmeter. When this butterfly valve

bedisc was perpendicular to the flow, the entire cross sectional area of the

supply line was closed off. However, for frequencies of rotation between

20 and 69 cycles per second, the area was closed off for such a small

portion of the period that the pressure downstream in the model plenum

never varied more than 15 percent from the mean plenum pressure. Asdiscussed in the text, the sinusoidal plenum pressure variation was

31

._ , , :. .-. , . " ' • . . .. i ,•, !,•,• ,•• ' ; i•" .•.' . . . . •. .

recorded by oscillograph traces and the mean was used to calculate jet

velocity. Upstream and differential pressure across the venturimeter

were sensed by transducers whose output was displayed on TIC readout units,

where the mean was recorded by hand. When these venturimeter readings

proved questionable, the isentropic mass flow was calculated using the

previous equations and the calibrated variation of slot height with duct

pressure as shown below. L

.029 -ý-h .027 + .0001 Pd[I

.028-

h, inches.027 LI.026L .U- .... ..

0 5 10 15 20 LI

Pd, in. Hg. ii

* a ~~32~- _

¼

S• ' • qd • • " r, : . q ' d Or .I

UTABLE - I GEOMETRIC PROPERITES OF NACA 66-210/CCW

AIRFOIL SECTIONS

Parameters common to all four configurations:

c' - 7.083" c " 8.000"

6f - 180O

r - 0.276"

h - 0.009"

r/c' - 0.03897 r/c - 0.0345

h/r - 0.0326

h/c' - 0.00127 h/c - .00113

Xglot M 6.819"

Xslot/c' - 0.9627 Xs1 0 t/C - 0.8524

t - 0.813"

St/c' -O0.11 48 tic - 0.0102

rLE 0.0530"

HrLEc' - 0.00748 rLE/c - 0.00662

Xdroop = 1.200"

Xdroop/C' - .1694 xdroop/c = .1500

* LI Leading Edge Deflection:

Configuration

CcW 6LE, deg Cdroop/c cdroop/e'

800 0 .8854 1.00000

815 15.3 .8879 1.00282

829 29.0 .8853 0.99989

844 43.6 .8711 0.98385

(cdroop is chord from drooped leading edge to rotated trailing edge,

but is not used in coefficient calculation)

S"33

I________ LI

- -

U,a a a a

C 0

N U0 0

-

�ju * .�.- HM V� *� 4% M i 0U .4 S U% N .4 -4 �4 N -

- .4 .4 �N 482 8 8 8�8 S [IU

-t 0 0 0 0 0

0' . .

4 LI* N

- �0J -4 Ii� � � 0% 0%

-4 Ps 4 e� NO * N N

______ Sj� �U -4 -4S*��*�Sj Ii.

N* I.

44% Q'0 0�N 10 0 60-4________ UN *

5 0 44%-4 -4 0 -4

___- - Ko 0

�- e 4'.. N N Ps 44 NN0 N N N - U� NA� N 0000 0 000 0o *� * * . . .

I 0 0 0 0 0 0 0 0 C' 0

C4 ____o - - - ____ - IiU

N N 44%N

'4 _____ _____ Li-4 - -

U N N NN N 44% N3 N S N

4 * -4 -4 4 0 -4 1 �IN N N N N N

UU - - -U *� ___ __ LI

S'.4

a � a Ii*0

* r.A. I hi

* hiI � .� *� � -�

U Nb.

A _____L!. ___ ____

I7-

LIH

Wi, U

4j --

�iI I'.4(.

'4

aa -.-. 0 8

0 U '-4'.4 be"

LI t.4UU

LIu be.4J dJ

8oLi0

H -4

H I �

H 0 '-4IH I I

S

H

4 ii 35

.1

0U VulK

ILI

v4-400L00 "A

usJ

D0

36U770

-z77 -7,

i •I iI

S 'II

I 0

I.i - I

VA/ /w1- I

iI! -

tI.

L ' ~ N

77 -A

Ki

Q3.

V 4 A

--- it

§ U

.44

", .4S 4 -

04.

'I ! .I X

38

S.. 0'U

L V K ,, -.

1-ýý,W.4

L J4

-444

C- INC1 -4-ell

/c0k

00

W,4

u1 0

oul4

39 '

'---

-------

~~W3

UA

El C4

404

* � *'v- �

-r �

I*I

SK I I

* IiK I IiK I* I �I4

U

U��1

I.- II 'U.'

-I

II

I41

.1 I** We: .-. **..***,I,1*..*&

......................................

IIIIIII

IIIII

I, ft

IIII Figure 8 - Aft View Showing Wake Rake, Model, and Tangential Wail Blowing

I* 42

N - un . - *1

5.14 +.o3.3 x

3+ 0

1:o-

-6.9 zIa

5.

8.4

t_,11 3*S.

LID

COC

V.040 .080 .120 .150 .200 .4MOMENTUM COEFFICIENT, CU

Figure 9 -Lift as a Function of Blowing f or the NACA 66-210/CCW

Airfoils, q -30 pef, Re a 0.64 x 106, h -0.009"

(a) CCW 800, 8jE -0

43

L- .9

Cc, des U

53 U

52.3°

2 3*

CMio; - •5. 3"

CI" 8. 4*

Li•, •11. 5"

I.A

Flagged symbols: q - 20 pst i

CD'0 .040 .080 .120 .1o0 .2no .240

MOMENTUM COEFFICIENT. cl

Figure 9 (Continued)

(b) CCW 115, 6LE - 15.3"

4i

lTt"

a ctc deg ac

+. .

i- x 3. 38..2°1.2 +

I.C -.9 x 69--- 3.9 Z,16.3 0

CL

HYLJ =.....=_... -. =--• 14.5 o

- 16. 3"

K 9 Flagged symbols: q 20 psf

iN; I l

'0.010 .080 .120 .160 .200 .2'N0

MOtIENTUM COEFFICIENT. Cu

Figure 9- (Continued)

(c) CCW 829, 6LE - 29"

i45

IL r

ac. des SC " 5.2a U

C 5.2

14-4q 0D 3.2" *

1.2 ,,.!--s; 1.2%9

o +

II

1.21

CII

Flagged symbols: q -20 pof

o,.I I I I

0 0q0 .080 .120 .180 .200 ,240

OPMENTUH COEFFICIENT. Cu

Figure 9 - (Concluded)

(d) CCW 844, 6LE 43.6*

46 {

it4.0

3.0

6 .4*

CO1 & 0. 11

2. A I - I A . a . I

a d 21f 30 5d0

-12

LZ CU C1 cog% am

t- -, o .10•- 2.065 -. 335 259

++-10 •--• 15S3 .1064 2. $08 -. 3?7 751

-- - -, 29.0S .1055 3.380 -. 593 412

0- - 43.6' .1052 3.590 -. 604 542

(Symbol& Identify leadigl "nd t•rauaI- 8 edge points before droop)

2 0

L3 -6

0

+ 21

+21 • rop Hne PInt * +

0 .2 .4 .6 .8 1.0

X/Ce

Figure 10 - Effect of Droop on Preventing Leading Edge Separation,

NACA 66-210/CCW Airfoil, ac - 8.40, Cu" 0.11,

Experimental Pressure Distributions

47.. .

•16 1

LI- Curve C11 CJ c s Cd

0 A .000 -. 204 -. 01% 101

A I .016 .597 -. 205 .047

9 C .054 1.896 -. 534 -. 013-12 0 D .083 2.6'1 -. 712 -. 033

1 1 .127 3.532 -. 954 -. 053Upper Surface

-10 Lower Surface

Symbols Identify leading and trailiCp edge points before droop.

-8 I!

6U-' L LI

-4-? Il

0B C

Droqp Hinge Point Slot

0 .2 .4 .6 .8 1.0 HX/Co

Figure 11 - Effect of Blowing on Experimental Pressure Distributions for

NACA 66-210/CCW Airfoil, SLE = 43.6%, ac -0.860

48 --

7,-...

TIt1 a C C CM25 Run

rL -- • -6.91 .0885 2.230 -. 669 2460----- --U -3.GO ,0873 2.531 -. 637 226

[---- -0.68 .0855 2.488 -.499 154

-- 3.41 .0856 2.262 -. 347 173!,-0 -- -•-O 11.29 .0905 1.658 -.. 316 272(Symbols identify leading and trailing edge

points)

Li See Figure 14(a) CCW800-8

¶ I;p -6

- I2• -• . ..........

-691\,-3.80" 0 6

Slot+ 2"-- -- A I ,

0 0.2 0.4 0.6 0.8 1.0

X/C?

Figure 12 - Effect of Incidence on NACA 66-210/CCW with No Leading Edge

Droop at Cz • .088, Experimental Pressure Distributions] LI 49

-20

-18

-16 I

CIC CM8 Ct Cool Ito

-14 O-O) -6.90 .1721 3.S49 -1

Z.- - -- 0 1.23 .1700 4.285 -1.021 29S

h • ---- 3.11 .1741 4.367 -1.002 316

0- -0 6.43 .1727 3.692 -0.4,2 419]

-12 .-- '.-- 11.53 .1733 2.923 -0.364 "4 2

,Sysmloos ida.lt~y lodimn cud irailJ.1s

#d4e po.sta vefore 4dOOp)

see viture 14(c) CCI 629

V-10 Il

'C, I LI

' i!-6

Ol Droopstat Point •lot

00.2 0.,4 0.6 0.8 1.0OL

X/C°

Figure 13 -Effect of i~nci•dence on NACA 66-210/CCW, Ailrfoil- vitth •

29' Leading Edge Droop at C.Qs 0.173, Experimental

Pressure Distributionsi L

50 L

9, 4`_

SS

C11t

.12 1

LI

ac. deg

Figure 14 -Lift Curves at Constant Momentum Coefficient for

66-210/CCW Airfoils, q -30 paf, Re 06 0.64 x 106,

h . 009"

(a) CCW4 800. 6E 09

51

Li 7 ~ - -

.16I

.12.10de I

C.0

0.0

-8 --

Figure 14 - Continued

(b) CCW S15* SLE 15.3*

52U

.244

'---I

.24

.161

35

o4±.02

.. .- ComS tlIml NACA 66-210

N. 3 ao106)

Gt %T Cu - .141

.- .56 4 x 10

*-O" split flI

Sf •--'-. -- Clean

Figure 14 - Concluded

(d) CCW 844. 6• - 43.6" *

54

iLi..1 6"" .. ! - .. : . - . .• . ... . ,• • - i•

LIR .64 3. 106

h .009"

6 -43.6

CCx 29 .24

15 3K .20

5Li -g - .0

". 4 -. 12

Flap '

2 .

Conventional 66-210'

1

Re -3 x 106

n _ t I1I , iI I_

I... I I •

-8 -4 0 4 8 12 16

•STALL, deg,

Figure 15 - CLmax ard OSTALL as Functions of MomentumCoefficient and Nose Droop Angle for

66-210/CL14 Airfoils

LI ~7777777T7755

L:]i

6

Re A 0.64x 106 8L e 3.

h .009" . g:

5 -. , --

[1

iI3

2

0 0-.04 .0S .12 .16 .20 .24

CUIli(r iC - as; a Function of Droop and Mmentum

Coefficient, MIACA 66-2lO0CC'=- Airfoils

56 I

1.4 E

1.2 q -10 s

11.0

iis 20 26 30' 36 41 4.23

31 4

21 -I a-0. 8

0 _I.

0. 0.6 .7 0.

0.9Re3 t 10-6

Pigur* 17 -Effects o1 Reynolds NuaberonLf

66- 2 10,#ccw Airfoil, CCW 800, Coeff0cient for009"Li 57

77 _

S. 4* Ik

i ~/X

-3.90

N [J

Cd

' ~-6.9*

8.11 A5.,1 + "'° )II-3.9 x

-6.9 Z

,0 .040 .060 •20 .160 .200 -. 2401WOENTUM CMFll: C IENT. c. !

Figure 18 - Drag as a Function of Blowing for the NACA 66-210/Ccw

Airfoils, q " 30 psf, Re " 0.64 x 106, h" 0.009"

(a) CCW 800, SLE 0*

58

- i

Li *

L1

L *Li

Cd

0 .1 it

-e.g x 71.aga~ symbols& q - 20 Paf

~~ IIOWINTUW COEFFICIENT. CU

"11"ure 18 (Continued)(b) CCW 815, 8L - 15.3*

!I

K.i........

icc. a .. 1'1111

4 .,*

S--0.

3.2 6 32

pligled symbol&: q 2C pef

i,2i

I LI

""0 .O .00 .120 .160 .200 .240

MM*WMTU• COFFICIENT.

Figure 18 - (Continued)

(t) CCW 829, *LE 29*

60

Li d

14i G 4.

it-4 A

-.. x

Floggd sy b W 0 s

-. 11 1 : .080 CcWa20 u q.1 0 .200

"MIiC FCEK.CFiur 18 (ocldd

(d)ZA ~ L IC 4,8 36

Se. d•q

3.3 x

11.31 . 3 4

ct)t

c a 1.3V

o 54_L

'1. 330

-6.9" -3.9

0 -:

"0 .N .0 5 • .12 • .18 ,20 24N

HqO WE TU M COIEIF FICIENT. C U l

Fi gure 19 - Pitching Mo mnt a n a Funct io n of Blowing for the NACA 66-210/CC W

Airfoils, q 30 pf, Re w 0.64 x 106, h 0.009"

(a) CCW 8O 0F CIp . 0"'62

Figure~~~~~~~~~~~~~ 19- -icigM7eta ucto7fBoin o7h AC 620c

Li

L i a c 1 1 . 5 0

5.30

2.300.80*

.. -3.9 °

-6.96

i j i

N De .O ,.8•12

.16 .20 .24! MOMENTUM COEFFICIENT. Co

1i Figure 19 -(Continued)L(b)

CCW 815, 6LE 15.30

Li

S 63

N P.I . . , ' " "

V4

aci14.5B

S8. 4

2.

1.2 S.2 31*20.1

-igur 19Z(otne)L

-649

U --. *

to .0 .08.12 18 .0 .2

Lii1 4.40

i 2 5

Li~ .4.

Lc c, deg ',.3

Li-0 . D 8.35.2 X3.2 *

Li :-3.9 Z

.20 3.2'A

MOMIENTUM4 COEFFICIENT. CU~

Figure 19 -(Concluded)

(d) CCW 844, 6L - 43.6*

Li 65

Me degU Ito12.1 A 9.0

9.0+

6.0 x 6 03.0 Lb

.0 +-6.0 3. 0

Flge sybos 30 2u, ~19 0, h -0.27

(a) CCW 241, .f - 180% 6LE - 30*

66

CO3. 0*

6.06

U 12.1

U

6.0 x"3.0

U

II I I,

IMO M•I N1IUM COEFIFICI D4T . cl•

Figure 20 - (Continued)

(b) CCW 242, 6f - 900, 6LE - 30"

67

t ~~~~ ~ ~ NI A * -. ' a~ t

-- t;: "°

C -" " 6.0'--- 3.0 .

I-A~. A .0 0.09U

.0e"/J .- A

.1. "'0 10 , _rl

C! 1-0

// cc, des15.2 (D

C~t12,1•"9.1 .4. Li6.0 X3.0 +-3.0 X

, -6.0 Z

LI ,,

CD

I. I tI i fj,0.040 .080 .120 • 160 .200 -240 ..

MOMENTUM COEFFICIENT. C1

Figure 20- (Continued)

(c) CCW 243, Extended T.E., 6 LE * 30*

LL

• , . ,; -,•,6 8.. U:@ • • .. • , 7••, ! ,.. , .•• ,: .. ..

S--[I lI I I I I -- I I

IiC!0 S c 6.06

/II

9 °*

0.00

-6 o

15.0 +

S6.0 4ELi Re-. 0

9 6

12. P

.(40 .000 .120 .160 .200 .-NOL! H~~OMENTUMq COEFFIC (IENT. CU

!! Figure 20 -(Continued)

(d) CC'W 2441, Reduced T.E., 6LE -30"

cles do

3.0 +~.0 x 3.0i

-6.iiC -6.0

ini2 .0

yo.040 .060 .120 .160 .200 .210HOMDTUI4 CMFPCI9I#T. C~

Figure 20 -(Concluded)

(e) CCW 245, 6f 180% 6LE 0*

70

241 I39

ase z 106 1.98 0.64

30 29*

L 21.27" 7.06".]r .875 .276

h .027 .009

Sh/c' .00127 .00127

6 r /' .0412 .0390

Lit/c' .1355 .1148

-x/c)slot .960 .963

SI h/r .0309 .0326 CCW 2415

4 Ol CCw 829

i, 7

o*I , I ~ I I ¼

0.04 .0O8 .12 .16 .20

CU

Fitlur 21 - Small and Larpe Chord CCW Airfoil Comparisont

C18

4- 241 .0611 0.0 .49 3,01W--- 557 244 .0110 0.0 .238 3.00

--- 56 141 .01 ,108 ).1 1.53 ',U0

6-- --- 5 144 .0411 .103 4.053 1.03-12 (3-1 degi eda cde1+

Edge 1• te /dkl eg DI e 1 ,)1dil l I;//•dle Poiant Wlelter Oro"I) +

CoI

-10

I III

-2 h

a Point Slot

+ [I÷ 2 + , ., I ,. . + .. . . I

0 .2 .4 ,6 ., 1.0 La/c

Figure 22 - Effects of Redaced Trailing Edge Radius - Comparison Between

Configuration CCW 241 and CCW 244 Experimental Pressure

Distributions 11

(a) accz 3m L

72

m• . • , •5+ .• - ., ., + • + , .+, ;. . ,, ,-+,.

Run Confis rc' C CjtLi0 0--- -0 59 241 .0411 0.0 1.493 12.11O- 594 244 .0210 0.0 1.121 12.09

-. -i 95 241 .0411 0.045 3.176 11.65

- "• 599 244 .0210 0.045 2.858 11.61

(SymboYs Identify Leading and Trailing EdgeI Points Before Droop)

-6

CpI4 S•I I••I

Run, 9 5

2" '" 5 9 9

- ~~~594 , .

DrooI Hinge Polat lot

0 .2 .4 .6 .8 1.0

011 KX/C

' Figure 22 - Continuad

(b) mc - 12*

73I IA

%TAI

Iiip

.08 a00Sf 00

-8 02 * 16001

Cov.i2a04-12 .0Spi.0a0,e-6 010*1 0-I

Fiue.16 itCre tCosatMmnu ofiintfr6A22CWArol

(a) CC .00,8fl~,~

30 p-4 4 1.9 12 101h607

i, L

12 .08C]0

H .02

* Li 16

.. 0

0. ac or aefft deg

Figure 23 Continued

M-7CC 242, Af 90, 6,E 3d

I 75

ac VS. CAa off Iii C

owll

- [11

9. 12.16 -

/ LXIFigur 23 Cntn

(c) - .C 243 LxeddI.. S

/ ~'~2 t76

off Ve C

*SAL~6

- -l

.12

Li .08

.04-

-8 4

a or ~eff. dog

Li Figure 23 -Contin~ued

(d) CCW 244, Reduced T.E., 8LE -30

77

344

TALL

5U- 16

L Li

-4 08 12 16 L0.0 a or eff.deR

-1

Figure 23 -Concluded [()CCW 245, 6f 180*, 0'

78 -?I

i4� �2' � u�-� - -, s.-.

iii

0(I

6 0U U

U -4 '05 -. 4

A a4.4U CLi ____________ -4 .6.4I- / 'I

/ 4,o,..' N

N -- -A 'a.4Li �4I

'.4Li 0

52C* o

*6.4'aU)0

-?Ii * 0II I.II

N 0U

ii'I"H .C

( -4 4-4Li N 0

'"4'-a

'q

".4

64.4U

U�j �6*

'U, A

Li>1

LI[1 79

i*. --- '- -

I

CCW 243 (CM0.16 1

ILIA%

4" 0

CC 245

2

I.8.-4 0I4-8 12 1 6

Figure 25 -Lift Curve Com~parison for the 64A-212/CCW Airfoils At

cu 0.16, mg s. 15', h -. 027", q -30 paf

so

1K....A- I& KA................................................

UC 4-

Li CCil 245

L1 max 4

Li2Z1 30 180 0.875 .0411 .0365242 30 90 .875 .0411 .0365243 30 180 .875 .0353 .0365 Extended T.E.

J, 244 30 180 .438 .0210 .0182 Reduced T.E.1245 0 180 .875 .0411 .0365

Li ~0

0.04 .08 .12 .16 .20 .24

CI a

Li Figure 26 -Maximum Lift Coefficients withini Test Limitations for5 Configurations of a C. C. Wing on a 2-1) NACA

64A212 Airfoil (m 1 150, h *.027", q *30 pef)

81

te7l

.04

L '1 ' Ii

.12,.o24A 1(3i

.16 .6 LI

.0.

P2B

"~~.0 ---."02 V1

y .02 0"0[i

-8 -4 0 4 S12 16

Figure 27 eia nd Stall Angles for the 64A-212/CCW

32

i _ .. . . ... - -- ' -4 A""7: .•-• _• •,.*• .F 7 7 7 7 -7 , :: , ,• : . ,- : ,• , .. -• -

4 10 psi

.2 0 h / .027" 0

.20 p0f

h " .027"0- 30 pf

i6

h - .027".

SCU q a 30 pof

CM .14: g"

*ee q" q 30psf1 , 00,/ de3o~

.12H~ a .027"

qa SO pof

.10

0 h .014" W

Li fl,

do.04 ... Y. -. qP30 psf.06

S- Choked , 14.8 pala

2.5 275 30 3.Pd/p.

0 1.25 1.5 1..75 2 .0 22? 2 2

07

0 10 20 30 40 50 60 70

Pd, in. IS.

Pigure 28 - Slot Height and Dynamic Pressure Effects on Cu,

CCW V41, ag 9". c a24"0

I' ~83 '

S, • E,•,..:••,i.';,.;:• • •• • . ... :: • :: ::::. •, , • •* , *, :•,:,,, .. ,, ,77•.

6.0 [

h - .020" h .021"5.0 --

4.5

3.0.

2.5 . 1

0. - i .85 .00 029 .09

£h -.014" .10 .06

1.5 pL/- 2 1.2 ". h -- 0 . 027 .004 12

1.1

1.0

0 .02 .04 .06 us: .0 .12 .14 .16

Cu

Figure 29 Slot Height and Pressure Ratio Effects on CCW 241,

ac A 90

84 ,.

~ ~ A~k 4s~

Li.05.0

ii 1 4~~~.5 / in0 O

3.0

2.5

U 2.0

1.5h.i. /rh I

1.i~.014 .0320 .00067G .027 .0517 .00130

1.0 c .054 .1234 .00259

0.5

0

-0.2 .04 .08 .12 .16 .20 .24

LilFigure 30 -Slot Height Effects on CCW 244, cc 0*

11ý

.33 ~CIO.

3

t/2.3 D

[ / 2 __

02

.50 0

30 40 AU

1.0 1.2 1.4 1.6 1.0 2.0 2.2 2.4Ite i 10-

Figure 31 - iffect Of RoYaolds Number on Lift, CCW 245(6LP 00, af 180 ac h * 027"

8[p*t.qý ~

.

-!- - - -w

"12.1 a$.0 4.6.0 x-. 0*

Ia - 15.1o

Cd

3.0°.

""* 0 i

'0.-a QOW .120 .160 .200 .240

1@KNTUM COEFFICIENT, CU

Figure 32 - Drag as a Function of Bloving for NACA 64A-212/CCW

Airfoils, q - 30 psf, Re = 1.98 x 106, h 0.027"

(a) CCW 241, 6 f =180%, 6LE - 0

87

- 7i1*~

LiI- a.e. 12.1. [A

1 9.1

1510

3.1 +

.0 *

U -.. 0

Cd 4.--.....

-6.0°c " "

t •,•- .0* ":

6, 0'

S I . . . I i I

"10 .040 .OO .120 .160 .20D .240

MOIENTMO CME]ICINIT. c.

Figure 32 - (Continued)

(b) CCW 242, 6f - 90", 6LE 30*

88

pb

IS.2 40

G€ 1.S.2" 1201 &.12 +

6.0 X-.0 *-3.0 R-5.0 Z

12..: a

i+ I ,

'0.d4 .1.06 .10 .5 "0 2

66. ENU~ 0*F11NT

3.0- ~-6.00

3.• , 0.0"

1 0 .0:NO .000 • 120 • 160 .200 .NO

"OMETU" COEFFIC~lel. C.

Figure 32 - (Continued)

(c) CCW 243, Extended T.E., 6LE - 30O

89

S.O 40

is .o 012.0 A9.0 4.6.0 x2.9 00o -. 0 4.

12.0 0o /

• i!

Cd LnL.

A 9.00

Q1100.0"

-6.0*x 2.9"

ol

Ln.

- I I * I I

1'0 .040 .080 .120 .160 .200 .240MOMENTUM COEFFICIENT. C,

Figure 32 - (Continued)

(d) CCW 244, Reduced T.E., 6 LE 30°

90.

- -i/w / -Ii

3.0..t. -6.0" %S6.0°.

8+

I -0.00

C

I S'+|

L!

S__ __ _O I.Ohi .0 .I120 .160 .200.2O

ENTIn6 cEFF1c1o•T. C,

Figure 32 - (Concluded)

(e) CaiY 245, •f - 160", 5E - O

91

....... ( 0o..133)

.02

.01

Cd

COJ 243

0-.0.1201 2

.02 -45 CCUa 241 5

!I-. 034

Figure 33 - Drag Coefficient at C wMax Within Test Limits for NACA

64A-212/CCW Airfoils.

ag _ 15", h - 0.027". q =30 psf

92

*'' del

12.1 A

I S•

U2 5 ' ac 15. 10

C3 12.10

4 L-6.0, 6

"1.003.0*

0 .04 .0 -12 •le .20 .214

POINENTIM COFFICIENT. c.

Figure 34 - Pitching Moment Variation with Blowing for CCW 241,

h - 0.027", q - 30 psf

j 93

I; ~ '~-

3 C13

C.,

ICI

-. .4-. -1.4 ~

for MACA 6AA-212iCC-7 Airfoils (h - .027". q 330 pf)()cWd 241, 6f l dOo, 56 - 309

4 ~94~

[1

L

I

I'I

II

9.A /.0. 12

5

' . .X12. 1_.

'y-J-. 1

. (•, c• .02

S-.2 -. 4 -. 8 . -1.0 -1.2 -1.4 -1.6•

IFigure 35 Continued

(b) CCW 242, df = 90", S " 30'

195 95

!,I

........._n____

-I___-____ __....

" ~Uj

Li

LI I

3

g " •.08

.06

f z/• Cu .02

-. 2 -. 4 -. 6 -. 8 -1.0 -1.2

Figure 35 - ontinued

(c) CC• 243, Extended T.E.. 6LE- 30*

S~96

LI•L

CL 3

4 0

I 3I1F.1

I'2K.06

-. 2 -. 4 -,6 .8-1.0 -1.2

Figure 35 -Continued

(d) CCW 244, Reduced T.E., SLE 300

I ~¶ 97

[1.

3U.0I82H a -

; .o016

0

r .2 -. 4 -. 6 -. 8 -1.0 -1.2

Figure 35S Concluded(e) CCW 245, 6f- 180%, 6LE- 0*iJr-

9i

Li

S.04 .OS .12 .16 .20

H0

-. 2

I -. 4

I -. 6

I! - .8 "

CCW 241

1K-1.0

CCW CCV 245

-1.. 2 243 CC 244

I -1.4

.-1.6 • cc 242

Figure 36 - Pitching Moment Coefficient at Cimax within Test Limits

for NACA 64A-212/CCW Airfoils, a.g _ 15,

h 0.027", q - 30 psf

99I _ _I J -

3.5-.

30 -,

a, pulsO/sec

02003.0-

50

0 20

0 60

r69

2.0-

1.5 -

1.o ./ • • -0o, q. -30 pe,, h.*.027"

0.5

.02 .04 .06 .08 .10 .12 .14

C, f�rou venturimeterFisure 37 - tesn Lift and Momentum Coefficients as Functions of Pulsing

Rate, CCW 244

100 1

. 028w, 40

-20

.024

40.020 6

.016

,012: W. pulses/sec

F1 30 q -30 psf

t _ _ _

L A 50 h -. 027"1

.004 Q 60Di 69

0 4 8 12 16 20 24 28]Pv, psi$

(line pressure Into vouturl)

Figure 38 - Mean Mass Flow Measured by the Venturimeter, CCW 244

101

3.5-

W. pulses/soe0 0 .,, 0 eqs0 20J 30

3.0 4. 50O 60r1 69

2.5

0]

2.0

1.5

1.0

S6-00q 30 pot

h , .027"

0.5

0

.02 .04 .06 .08 .10

-0. 25 - (expanded to Q,)

Figure 39 - Corrected )ean Lift and ,omentum Coefficients as

Functions of Pulse Rate, CCW 244

102

6

- 0 Clean Airfoil -Clean

5-• Full Spoiler Airfoil 6 " -

--- Slotted Spol lero

4

CL

2 2j.~ =5. 5

51

-I~~ ~~~ 18 , 0

0 .04 .08 .12 .16 .20

CU

Figure 40 - Effect of Spoiler Deflection on Lift of Elliptical

C.C. Airfoil (ag - 0, q = 20 psf, h - .009")

103

,0 I•20 / 50

-1 . - -• .06 t _1_•

""Ct_2___ _ ..... __ _ LI

-3 I-2 i

A Ci - Clclean fS0°, q 20 psf, h .009"1

- -- Full Spoiler

--- ,&-- -- Slotted Spoller

Figure 41 - Lift Lose as a Function of Spoiler Deflectto

104 •

.. H-

-I

1 -..- -

.0 .. 4 ... 4,5,

-. .. f

-BII 4-4

9!

I05

5 4- -.* -

Clean AifeoLl so |1°.35--- Fell Spoiler SI

- - - - - Slotted Spoiler

/[j

,01.30

.0 .0-,I I

S~clean Airfoil

as = 0% q w 20 per, h .009"

Cd gure 43 Drag Increase due to Spoiler DeflecI.ion

on Elliptical C.C. Airfoil

/i

SUU

1.00

'A=0 -2 a, 09

25.f' 4c,,.1 .04 .08 .12 .16 .20

-. 2

-. 4

"'25 -. 2-2..

-1.4 L 0 p fClean Airfoil_-h 009"

Figure 44 -Pitching Memmut Variation with Spoiler DeflectionL107

F.