NASA 

CR 

3523 

. c.1 

c -1 

NASA Contractor Report 3523 

Wind Tunnel Tests of High-Lift 

Sys 

for Advanced Transports Using 

High-Aspect-Ratio Supercritical wings 

John B. Allen, Wayne R. Oliver, 

and Lee A. Spacht 

CONTRACT 

JULY 1982 

NASl-15327 
TECH LIBRARY KAFB, NM 

Ollb2217 

NASA Contractor Report 3523 

Wind Tunnel Tests of High-Lift Systems 

for Advanced Transports Using 

High-Aspect-Ratio Supercritical Wings 

John B. Allen, Wayne R. Oliver, 

and Lee A. Spacht 

McDonnell Doughs Corporation 

Long Beach, California 

Prepared for 

Langley Research Center 

under Contract NASl-15327 

National Aeronautics 

and Space Administration 

Scientific and Technical 

Information Office 

1982 
t 
CONTENTS 

SUMMARY ................................ 

FOREWORD ............................... 

INTRODUCTION ............................. 

SYMBOLS ................................ 

PART I. WIDE BODY DC-X-200-TYPE MODEL ................. 

LB-486B,C MODEL DESCRIPTION ...................... 

LB-486B,C INSTRUMENTATION ....................... 

LB-486B,C MODEL INSTALLATI'ON ..................... 

REVIEW OF PHASE I RESULTS ....................... 

LB-486B,C RESULTS AND DISCUSSIONS ................... 

Reduced VCK Deflection ...................... 

SealedSlats ........................... 

Fixed-Camber Krueger ....................... 

Slat Trim Effects ........................ 

High-Speed Aileron ........................ 

Two-Segment Flaperon Replacement ................. 

Differential Flap Deflection ................... 

Ames 12-Foot and Langley V/STOL Tunnel Comparison ........ 

PART II. NARROW BODY ATMR-TYPE MODEL ................. 

LB-507A MODEL DESCRIPTION ....................... 

LB-507A INSTRUMENTATION ........................ 

LB-507A MODEL INSTALLATION ...................... 

LB-507A RESULTS AND DISCUSSIONS .................... 

Cruise Wing Characteristics ................... 

Reynolds Number and Mach Number Effects ........... 

Nacelle/Pylon/Strake Effects ................. 

Landing Configuration Characteristics .............. 

Reynolds Number and Mach Number Effects ........... 

Nacelle/Pylon/Strake Effects ................. 

Page 

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48 

48 

48 

51 

51 

53 

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61 

65 

67 

67 

72 

80 

93 

93 

101 

Large Inboard Flap Deflection . . . . . . . . . . . . . . . . 108 

iii 
CONTENTS - (Concluded) 

Page 

Takeoff Configuration Characteristics .............. 108 

Strake Effects ........................ 117 

Mach Number and Reynolds Number Effects ........... 117 

Alternative Flap Settings .................. 127 

Aileron and Spoiler Characteristics ............... 127 

Landing Gear Effects ....................... 135 

CONCLUSIONS AND RECOMMENDATIONS .................... 142 

Conclusions ........................... 142 

Recommendations ......................... 144 

APPENDIX A LB-486A,B,C, CONFIGURATION NOTATION ............ 146 

APPENDIX B LB-486A,B,C, DIMENSIONAL DATA ............... 150 

APPENDIX C LB-486A,B,C, SLAT GRID NOTATION .............. 154 

APPENDIX D LB-507A CONFIGURATION NOTATIONS .............. 164 

APPENDIX E LB-507A DIMENSIONAL DATA .................. 168 

APPENDIX F LB-507A SLAT GRID NOTATION ................. 172 

REFERENCES .............................. 175 

iv 
Figure 

ILLUSTRATIONS 

Page 

High-Lift Low-Speed Wind Tunnel Model ............ 15 

High-Lift Components Evaluated in Experimental 

Test Program ........................ 16 

Wing Diagram (W3B) ..................... 18 

Horizontal Stabilizer HIA Diagram .............. 19 

Vertical Stabilizer VIA Diagram ............... 19 

Nacelle/Pylon N2A P2A Diagram ................ 20 

LB-486 Leading Edge Device Gap, Overhang, and 

Deflection Definitions ................... 21 

LB-486 Flap Gap, Overhang, and Deflection Definitions .... 21 

Model Installation in the NASA Ames 12-Foot 

Pressure Wind Tunnel ..................... 23 

Model Installation in the NASA Langley V/STOL 

Wind Tunnel ........................ 24 

Tail-On Characteristics for the Cruise Wing With 

Nacelles, Pylons, and Strakes Attached ........... 25 

Effect of Reynolds Number on Clean-Wing Section 

Maximum Lift ........................ 26 

Tail-Gn and Tail-Off Aerodynamic Characteristics of the 

VCK With Two-Segment Takeoff Flaps Configuration ...... 28 

Tail-On and Tail-Off Aerodynamic Characteristics of the Slat 

With Two-Segment Takeoff Flaps Configuration ........ 29 

Effect of VCK Deflection 

A. Lift and Pitching Moment ................ 31 

B. Drag .......................... 32 

Effect of Sealed Slats 

A. Lift and Pitching Moment ................ 34 

B. Drag .......................... 35 

L/D Comparisons for Clean Wing, Sealed Slat, and Landing 

Slat Configurations ..................... 36 

Effect of Inboard Sealed Slat With Landing Flaps 

A. Lift and Pitching Moment ................ 37 

B. Drag .......................... 38 

I 1 

2 

8 

9 

10 

11 

12 

13 

14 

15 

16 

17 

18 

V 
ILLUSTRATIONS - (Continued) 

Figure 

19 

20 

21 

22 

23 

24 

25 

26 

27 

28 

29 

30 

31 

32 

33 

34 

35 

Effect of Leading Edge Device With Landing Flaps 

A. Lift and Pitching Moment . . . . . . . . . . . . . . . . 

B. Lift-Drag Ratios . . . . . . . . . . . . . . . . . . . . . 

Comparison of Full-Span FCK and FCK/Slat Combination 

A. Lift and Pitching Moment ................ 

B. Lift-Drag Ratio .................... 

Grid Study for Short-Chord FCK/Slat Combination 

A. Lift and Pitching ................... 

B. Drag ......................... 

Comparison of Full-Span Slat and Short-Chord FCK/ 

Slat Combination 

A. Lift and Pitching Moment ................ 

B. Drag .......................... 

Effect of Slat Trim With Landing Flaps 

A. Lift and Pitching Moment ................ 

B. Drag ......................... 

Ames 12-Foot and Langley V/STOL Comparison ......... 

LB-507A Model Three-View .................. 

LB-507A Wing (WIB) ..................... 

High-Lift and Lateral Control Surfaces ........... 

Leading Edge Device Gap, Overhang, and 

Deflection Definitions ................... 

Flap Gap, Overhang, and Deflection Definitions ....... 

Dorsal Fin (D2A) ...................... 

Horizontal Stabilizer (HID) ................ 

Vertical Stabilizer (VID) ................. 

Pressure Row Locations ................... 

Installation in NASA Ames 12-Foot Pressure Tunnel ..... 

Cruise Wing Characteristics and Repeatability 

A. Lift and Pitching Moment ................ 

B. Drag ......................... 

Page 

. 

40 

L 

41 

42 

43 

44 

45 

46 

47 

49 

50 

52 

56 

57 

58 

59 

60 

62 

63 

64 

66 

67 

68 

69 

vi 
ILLUSTRATIONS - (Continued) 

, 

Figure Page 

36 Comparison of LB-507A and LB-486B Cruise Wing Characteristics 

A. Lift and Pitching Moment . . . . . . . . . . . . . . . . . 70 

B. Drag........................... 71 

37 Mini-Tuft Photos for Cruise Wing/Body With Nacelles (Run 3) 

A. 'FRP =12.58' . . . . . . . . . . . . . . . . . . . . . . 73 

B. "FRP = 13.61' (clCL ) . . . . . . . . . . . . . . . . . . 73 

MAX 

C. "FRP -14.59O . . . . . . . . . . . . . . . . . . . . . . 74 

D. "FRP =16.54' . . . . . . . . . . . . . . . . . . . . . . 74 

E. 'FRP =18.49' ...................... 75 

F. "FRP ~20.44' ...................... 75 

38 Chordwise Pressure Distributions of Cruise Wing With Nacelles, 

Pylons, and Strakes Attached 

A. "FRP = 13.61' (aC 

'MAX 

) .................. 76 

B. CLFRP =14.59O ...................... 77 

C. "FRP =16.54' ...................... 78 

39 Effect of Reynolds Number on Maximum Lift of Cruise Wing . . . 79 

40 Effect of Mach Number on Cruise Wing 

A. Lift and Pitching Moment . . . . . . . . . . . . . . . . . 81 

B. Drag........................... 82 

41 Effect of Nacelles and Pylon on Cruise Wing 

A. Lift and Pitching Moment . . . . . . . . . . . . . . . . . 83 

B. Drag........................... 84 

42 Mini-Tuft Photos for Cruise Wing/Body (Run 113) 

A. "FRP =12.54' . . . . . . . . . . . . . . . . . . . . . . 85 

B. "FRP = 13.55' (aC 

LMAX 

) . . . . . . . . . . . . . . . . . . 85 

C. CIFRP =16.5'....................... 86 

D. "FRP =18.48' ...................... 86 

vii 
ILLUSTRATIONS - (Continued) 

Figure 

43 

44 

45 

46 

47 

48 

49 

50 

51 

Page 

Chordwise Pressure Disbributions of Cruise Wing Without 

Nacelles and Pylons 

A. "FRP = 11.55O (&L ) . . . . . . . . . . . . . . . . . 

MAX 

B. "FRP =13.55O ...................... 

C. "FRP =14.50° ...................... 

D. "FRP =16.50'.. .................... 

Effect of Strakes on Cruise 

A. Lift and Pitching Moment ................ 

B. Drag i ......................... 

Landing Slat Grid Optimization 

A. Lift and Pitching Moment ................ 

B. Drag .......................... 

FCK Optimization With Landing Flaps 

A. Lift and Pitching Moment ................ 

B. Drag .......................... 

FCK and Slat Comparison With Landing Flaps 

A. Lift and Pitching Moment ................ 

B. Drag .......................... 

Effect of Reynolds Number on Maximum Lift of Landing 

FCK/Slat Configuration . . . . . . . . . . . . . . . . . . . 

Effect of Mach Number on CLMAX Landing 

FCK/Slat Configuration . . . . . . . . . . . . . . . . . . . 

Effect of Nacelles and Pylons on Landing Slat Configuration 

A. Lift and Pitching Moment . . . . . . . . . . . . . . . . 

B. Drag . . . . . . . . . . . . . . . . . . . . . . . . . . 

Effect of Strakes on Landing Slat Configuration 

A. Lift and Pitching Moment . . . . . . . . . . . . . . . . 

B. Drag . . . . . . . . . . . . . . . . . . . . . . . . . . 

. 

87 

88 

89 

90 

91 

92 

94 

95 

96 

97 

98 

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100 

101 

102 

103 

104 

105 

. . . 

VIII 
Figure 

52 

ILLUSTRATIONS - (Continued) 

Page 

Effect of Strakes on Landing FCK Configuration 

A. Lift and Pitching Moment . . . . . . . . . . . . . . . . . 106 

B. Drag . . . . . . . . . . . . . . . . . . . . . . . . . . 107 

Effect of Large Flap Deflection 

A. Lift and Pitching Moment . . . . . . . . . . . . . . . . . 109 

B. Drag........................... 110 

Mini-Tuft Photos for 35' Flap Deflection Configurations 

A. O.H. = 1% (cxFRP = 19.12', 2' PAST aC 

LMAX 

) . . . . . . . . . 111 

B. O.H. = +l% (aFRP = 20.08', 4' PAST crCL ) . . . . . . . . 111 

MAX 

53 

54 

55 

56 

57 

58 

59 

60 

61 

62 

Effect of Large Flap Deflection on Trailing 

Edge Pressures ........................ 112 

Outboard Slat Grid Optimization With FCK Inboard 

A. Lift and Pitching Moment ................. 114 

B. Drag .......................... 115 

Mini-Tuft Photo of Takeoff Configuration Showing Effect 

of Outboard Sealed Slat 

A. Slotted Slat Outboard, ~1 = 20.94' ............ 117 

B. Sealed Slat Outboard, ~1 = 20.85' ............. 117 

Effect of Inboard Leading Edge Device With a Sealed 

Slat Outboard 

A. Lift and Pitching Moment ................. 118 

B. Drag .......................... 119 

Effect of Slats on Rolling and Yawing Moments Through 

Stall With Sideslip ..................... 120 

Effect of Inboard FCK slat Deflection on Ro.lling Moment 

Through Stall With Sideslip ................. 121 

Effect of Strakes with Takeoff Flaps and Sealed Slats 

A. Lift and Pitching Moment ................. 122 

B. Drag .......................... 123 

Effect of Mach Number on Takeoff FCK/Slat Configuration 

A. Lift and Pitching Moment ................. 124 

B. Drag .......................... 125 

ix 
ILLUSTRATIONS - (Concluded) 

Figure Page 

63 

64 

65 

66 

67 

68 

69 

70 

71 

72 

Effect of Reynolds Number on Maximum Lift of Takeoff 

FCK/Slat Configuration ................... 126 

Aerodynamic Characteristics of the 15'/10° 

Flap Configurations 

A. Lift and Pitching Moment ................ 128 

B. Drag .......................... 129 

Aerodynamic Characteristics of Clean Trailing 

Edge Configurations 

A. Lift and Pitching Moment ................ 130 

B. Drag .......................... 131 

Takeoff L/D Summary ..................... 132 

Rolling-Moment Coefficient Due to Aileron Deflection for 

Sealed Slat Takeoff Configuration .............. 133 

Rolling-Moment Coefficient Due to Aileron Deflection for 

the FCK/Slat Landing Configuration ............. 134 

Rolling-Moment Coefficient Due to Spoiler Deflection for 

the Sealed Slat Takeoff Configuration ............ 136 

Rolling-Moment Coefficient Due to Spoiler Deflection for 

the FCK/Slat Landing Configuration ............. 137 

Effect of Symmetrical Spoiler Deflection 

A. Lift and Pitching Moment ................ 138 

B. Drag .......................... 139 

Effect of Landing Gear 

A. Lift and Pitching Moment ................ 140 

B. Drag .......................... 141 
SUMMARY 

This report presents the results of the wind tunnel testing of an 

advanced-technology high-lift system for a wide body and a narrow body model 

of a fuel-efficient transport. These aircraft, derived from detailed system 

studies for a medium-range transport, incorporated high-aspect-ratio 

supercritical wings. Along with the wind tunnel results from an earlier 

phase of the program, these experimental results represent the first 

low-speed high-Reynolds-number wind tunnel data for such an advanced 

transport. Experimental data included the effects on the low-speed 

aerodynamic characteristics of slat, variable-camber Krueger (YCK), and 

fixed-camber Krueger (FCK) leading-edge devices, two-segment and 

single-segment trailing-edge flaps, nacelles, pylons, ailerons, spoilers, 

horizontal tail, and landing gear. Both Mach and Reynolds-number effects 

were also studied for selected configurations 

The cruise wings achieved tail-off maximum lift coefficients near 1.6 and 

tail-off lift-drag ratios near 21. For the high-lift configurations, the 

values of maximum lift coefficient were significantly improved when compared 

with current aircraft values. Typical tail-off maximum lift coefficients 

for takeoff and landing configurations were 2.4 and 3.1, respectively. 

Corresponding tail-off lift-drag ratios were 15.4 and 9.8. These ratios 

represent significant improvement over those of previous-generation 

aircraft. 

Aileron studies indicated that, for all flap settings, negative deflections 

(trailing edge up) were more effective than positive deflections (trailing 

edge down). The effect of spoiler deflection on roll characteristics 

indicated improved effectiveness as the flap deflection was increased. 

Symmetrical spoiler deflections, for both takeoff and landing flaps, showed 

the spoiler to be very effective in reducing lift and incremental .drag. The 

landing gear caused a slight reduction in maximum lift coefficient for the 

landing configuration. 
Analysis of the data has identified areas where continued efforts could 

result in further improvements. These areas include pitching moments for 

the high-lift configuration, and ground effect characteristics. Specific 

test items are suggested for this continued development. 
FOREWORD 

This document presents the results of a contract study performed for the 

National Aeronautics and Space Administration (NASA) by the Douglas Aircraft 

Company, McDonnell Douglas Corporation. This study was part of Phase II of 

the Energy Efficient Transport (EET) project of the Aircraft Energy 

Efficiency (ACEE) program. 

Acknowledgments for their support and guidance are given to the NASA 

technical monitor for the contract, T.G. Gainer of the Energy Efficient 

Transport Project Office at the Langley Research Center; to the NASA Project 

Manager, R.V. Hood, and to J.R. Tulinius, the on-site NASA representative; 

also to R.T. Whitcomb of the Langley Research Center for his concept of the 

supercritical wing. 

Acknowledgments are also given to the director and staff of the NASA Ames 

and Langley Research Centers, at which facilities the test programs were 

conducted. The cooperation and discussions concerning high-lift development 

for high-aspect-ratio supercritical wings of R.J. Margason and H.L. Morgan 

of the NASA Langley STAD Low-Speed Aerodynamics Branch were greatly 

appreciated. 

The Douglas personnel who made significant contributions to this work were: 

M. Klotzsche 

A.B. Taylor 

J.G. Callaghan 

D.E. Ellison 

W.R. Oliver 

G.G. Myers 

L.A. Spacht 

ACEE Program Manager 

EET Project Manager 

Branch Chief, Configuration Design and 

Development, Aerodynamics Subdivision 

Branch Chief, Stabilit.y, Control and Flying 

Qualities, Aerodynamics Subdivision 

Task Manager for initial-phase 

Aerodynamic Design (Report Author) 

Aerodynamic Design 

Aerodynamic Design (Report Author) 

3 
J.B. Allen 

J.D. Cadwell 

R.J. Commons 

F.B. Baugh 

D.H. Broderson 

L.B. Scherer 

G.K. Ige 

R.C. Leeds 

Task Manager for final-phase 

Aerodynamic Design (Report Author) 

Branch Chief, Aerodynamics Wind Tunnel 

Model Group 

Aerodynamics Wind Tunnel Model Group 

Aerodynamics Wind Tunnel Model Group 

Aerodynamics Wind Tunnel Model Group 

Aerodynamics Wind Tunnel Model Group 

Aerodynamics Wind Tunnel Model Group 

Aerodynamics Wind Tunnel Model Group 
INTRODUCTION 

The present investigation was made in connection with the high-lift studies 

of Reference 1 and the cruise performance studies of Reference 2. During 

the Reference 2 work, Douglas developed the high-aspect-ratio supercritical 

wing for the DC-X-ZOO, a 200-passenger wide body configuration proposed as a 

next generation transport. The results of that study were used to design a 

wing with minimum drag creep for the Advanced Technology Medium Range (ATMR) 

transport, a 176-passenger narrow body configuration. Both investigations 

showed that supercritical wing technology could significantly reduce fuel 

consumption and direct operating costs; they also established a sound 

technology base for future development work. 

The high-lift system reported in Reference 1 was developed for the 

DC-X-200. A model of a DC-X-200 with various leading- and trailing-edge 

high-lift devices was tested. The results indicated that although the 

system gave better performance than the high-lift systems on current 

transports, even greater improvements are to be gained by developing the 

system further. .Moreover, the takeoff and landing configurations tested had 

undesirable pitch-up at angles of attack near stall. Further investigation 

was needed to alleviate the pitch-up and improve the performance. 

The present investigation was undertaken to continue the high-lift 

development for the DC-X-200 (the effort reported in Reference 11, and to 

extend the development to the ATMR configuration with its narrower body and 

more advanced wing. Part I of this report describes the investigation of 

the DC-X-200 high-lift system. The same 4.7-percent scale model tested 

during the high-lift study was tested in the the present investigation, but 

with a number of the leading- and trailing-edge modifications that it was 

hoped would improve the performance. These included: 

1. A leading-edge fixed-camber Krueger which would be mechanically 

simpler than the variable-camber Krueger investigated in the work 

of Reference 1. 

2. A two-segment flap replacing the flaperon tested on the Reference 1 

model. 

5 
3. A variable-camber Krueger with a reduced deflection. 

4. A mixed leading-edge configuration (a slat outboard and a 

fixed-camber Krueger inboard). 

5. A two-piece leading-edge device, each piece having a different 

deflection. 

6. Changes in the slat trim, both next to the fuselage and around the 

engine pylons. 

7. A short-chord fixed-camber Krueger for the inboard wing to improve 

the pitching-moment characteristics. 

8. A sealed leading-edge slat to improve the takeoff lift-drag ratio. 

The model was tested in two different tunnels-- the NASA Langley Research 

Center V/STOL Tunnel in October and November 1979 and the NASA Ames Research 

Center 12-Foot Tunnel in July 1980. When tested in the Langley V/STOL 

Tunnel this model was designated the LB-486C; when tested in the Ames 

12-Foot Tunnel it was designated the LB-486B. These designations are used 

throughout this report. 

Part II of this report describes the ATMR investigation, in which the 

emphasis was placed on determining the effects of the narrow body 

configuration and the advanced wing geometry. These tests were made using a 

5.59-percent scale model (designated LB-507A) in the Ames 12-Foot Tunnel in 

January and Febuary 1981. The objective of the LB-507 program was to 

evaluate the low-speed aerodynamic characteristics of the narrow body model, 

including the following: 

1. The cruise wing characteristics. 

2. The influence of takeoff and landing slat configurations on the 

aerodynamic characteristics. 

3. Longitudinal stability characteristics (with and without the 

horizontal tail). 

4. Nacelle/pylon and landing gear effects. 

5. Spoiler and lateral control effectiveness. 

6. Mach and Reynolds number effects. 

7. Lateral-directional characteristics for selected configurations. 
The data obtained during the three tunnel tests included data on the 

six-component forces and moments. The data obtained in the NASA Ames 

12-Foot Tunnel included data on pressures measured at appropriate stations 

on the wing, slats, and flaps, and flow visualization photographs taken 

using a mini-tuft technique (Reference 3). The tests in the Langley V/STOL 

Tunnel were made at a Reynolds number of about 1.1~10~; those in the Ames 

12-Foot Tunnel at Reynolds numbers from 1.1~10~ to about 5.5~10~. 

SYMBOLS 

The longitudinal aerodynamic characteristics presented in this paper are 

referred to in the stability-axis system. Force data are reduced to 

coefficient form based on the trapezoidal wing area. All dimensional values 

are given in both International System of Units (SI) and U.S. Customary 

Units, the principal measurements and calculations using the latter. The 

model configuration notation is defined in the appendixes. 

Coefficients and symbols used herein are defined as follows: 

AR wing aspect ratio 

b wing span 

C wing chord 

cH horizontal stabilizer chord 

CD drag coefficient 

CL 

CLol = 0 

lift coefficient 

lift coefficient at 0' angle of attack 

CLMAX 

5 

(AC,) AB = -5’ 

maximum lift coefficient 

rolling-moment coefficient 

change in yawing-moment coefficient with a change 

in sideslip angle from 0' to -5O 

CM 

Cn 

(AC,) Af3 = -5' 

pitching moment coefficient 

yawing-moment coefficient 

change in yawing-moment coefficient with a change 

in sideslip angle from O" to -5O 

CP 

min 

minimum pressure coefficient 

CP 

TE 

pressure coefficient measured at the trailing 

edge of the element 

CV vertical stabilizer chord 

cW wing root chord 
FCK 

FRP 

HMAC 

iH 

(l-1 

L/D 

LH 

MAC 

MACH 

MS 

(RI 

O.H. 

RWMAC 

SH 

SREF 

SV 

SW 

TED 

TEU 

TS 

VCK 

VS 

WRP 

fixed-camber Krueger (flap) 

fuselage reference plane 

mean aerodynamic chord of the horizontal tail 

incidence angle between the horizontal tail and 

the fuselage reference plane, positive traili-ng 

edge down (deg) 

left wing panel 

lift-drag ratio 

distance between the 25-percent MAC point on the 

wing and the 25-percent MAC point on the 

horizontal tail 

distance between the 25-percent MAC point on the 

wing and the 25-percent MAC point on the 

vertical tail 

mean aerodynamic chord 

Mach number 

model station 

right wing panel 

overhang 

Reynolds number based on MAC 

horizontal tail area 

reference wing area 

vertical tail area 

wing area 

trailing edge down 

trailing edge up 

tunnel station 

variable camber Krueger (flap) 

Stalling Speed - the minimum steady flight speed 

at which the airplane is controllable 

wing reference plane 

. 

10 
. 

wuss 

x,y,z 

XH,yH 

xw 3 yw 

YVJV 

%lAX 

CIFRP 

clCL = 0 

AB 

r 

rH 

n 

GFAFT 

GFCK 

GFLAP 

6LE 

GSLAT 

GFMAIN 

6SP 

wing under slat surface 

spanwise, chordwise, and vertical fuselage stations, 

respectively 

spanwise and chordwise horizontal-tail stations, 

respectively 

spanwise and chordwise wing stations, respectively 

chordwise and vertical vertical-tail stations, 

respectively 

angle of attack at C 

+I AX 

angle of attack of the fuselage reference plane, 

positive nose up (deg) 

angle of attack for zero lift 

change in yaw (sideslip) angle 

dihedral angle 

horizontal-tail dihedral angle 

ratio of XN to semispan 

aft flap deflection angle, positive for trailing 

edge down (deg) 

flexible-camber Krueger flap deflection angle, 

positive for trailing edge down (deg) 

flap deflection angle, positive for trailing 

edge down (deg) 

general leading-edge device flap deflection 

angle, positive for trailing edge down (deg) 

leading-edge slat deflection angle, positive 

for trailing edge down (deg) 

main flap deflection angle, positive for 

trailing edge down (deg) 

spoiler deflection angle (symmetrical), negative 

for trailing edge up (deg) 

11 
GVCK 

A 

x 

variable-camber Krueger flap deflection angle, 

positive for trailing edge down (deg) 

sweep angle 

taper ratio 

12 
PART I 

WIDE BODY 

DC-X-200-TYPE MODEL 

LB486-C MODEL 

INSTALLED IN 

LANGLEY V/STOL TUNNEL 

13 

PART I 

WIDE BODY DC-X-200-TYPE MODEL 

LB-486B,C MODEL DESCRIPTION 

The wind tunnel model used for the program was a 4.7-percent representation 

of the DC-X-200 aircraft, and was the same as that used in Phase I of the 

EET Project study. The model is depicted in Figure 1. The configuration 

notation data, dimensional data, and grid position definitions are presented 

in Appendixes A, B, and C, respectively. The model was designed as a 

primary high-lift configuration that included a variable-camber Krueger 

(VCK). Secondary configurations employed either slats or fixed-camber 

Kruegers (FCK) along the leading edge. Combinations of an FCK inboard with 

a slat outboard were also tested. 

The primary trailing-edge configuration employed inboard and outboard 

two-segment flaps. Between these two flaps was a flaperon, essentially a 

single-slotted flap, that could be articulated in the same manner as the 

DIMENSIONS IN CENTIMETERS (INCHES) MODEL SCALE 

36.530 (14.382) I 

70.236 

(27.652) 

l- 

# 222.08 (87.435) I 

FIGURE 1. HIGH-LIFT LOW-SPEED WIND TUNNEL MODEL 

15 
main flap for the high-lift conditions, but that incorporated a high-speed, 

short-chord aileron in the retracted, or cruise, configuration. At the 

high-lift condition, this aileron was locked in an undeflected position. 

This permitted an 83-percent continuous flap span resulting in an improved 

span loading for high-lift conditions. The various high-lift components are 

depicted in Figure 2. 

SECTION A-A SECTION B-B 

[LEADING EDGE DEVICES/ (TRAILING EDGE DEVICES] 

PRIMARY CONFIGURATION - VCK PRIMARY CONFIGURATION - TWO-SEGMENT FLAP 

CLEAN TAKEOFF AND 

LANDING 

CLEAN 

TAKEOFF LANDING \ 

SECONDARY CONFIGURATION -SLAT SECONDARY CONFIGURATION -SINGLE-SEGMENT FLAP 

TAKEOFF LANDING 

SECONDARY CONFIGURATION- FCK 

FIGURE 2. HIGH-LIFT COMPONENTS EVALUATED IN EXPERIMENTAL TEST PROGRAM 

The model also included an aileron on the left wing panel, spoilers, and a 

remote-drive horizontal stabilizer deflection capability. Other model 

components included nacelles, pylons, landing gear, and a cruise wing 

trailing edge (i.e., flaps retracted). The fuselage consisted of DC-10 

model nose and aft fuselage shell sections, and a top center section and 

wing/fuselage fillet developed for Phase I testing. 

16 
A fuselage core was adapted for attachment of the fuselage shell sections, 

support of two !&module scanivalve systems, support of a bubble pack plate, 

and attachment of the wing and the vertical and horizontal stabilizers. A 

fuselage internal pitch system was installed in the core. This system 

permited the fuselage to be pitched from aFRP = O" to +lO" while the 

internal balance remained at oFRP = O". The other pitch angles were 

obtained by using the external pitch system. This system provided more 

accurate drag measurements between O" to 10'. 

The wing geometry and planform dimensions are shown on the wing diagram 

(Figure 3). The wing was designed to simulate the aircraft wing under a l-g 

load. It incorporated the following features: 

1. A cruise leading edge removable at the front spar. This leading 

edge was tested with and without simulated VCK stowage wells. Also 

provided was a WUSS (wing under slat surface) leading edge for the 

slat configuration. 

2. A VCK, FCK, and slat leading-edge flap device with variable 

deflection and position capability. 

3. A two-segment trailing-edge flap supported at five deflection 

angles by fixed brackets simulating the airplane flap linkage. 

Variable position capability was provided for the main flap. 

4. A manually set aileron, left side only, and spoilers both sides. 

5. Approximately 400 static pressure orifices installed in the VCK, 

slat, wing, and flaps. 

The geometry of the horizontal stabilizer is shown in Figure 4. The 

horizontal stabilizer was removable for testing tail-off. Each side of the 

stabilizer was fabricated in one piece without elevators. A remote control 

system was used to vary the stabilizer incidence between +5O and -15'. 

The vertical stabilizer planform is shown in Figure 5. The stabilizer was 

fabricated as one piece without rudders and was removable to provide a 

tail-off configuration. 

17 
40 

7) 

/ 

/ 

- 121.964 146.025l 

I%:..,,,, 1: 

DIMENSIONS IN CENTIMETERS (INCHES) MODEL SCALE 

lC/4)MAC Ill.3171 

Y = 160.280 163.1021 .I -. Fl.. ,.^ ^_A, 7 I 

xv4 = 36.367 

114.318) 

28.744 I / / 

FRONT SPAR PLANE 

/...-- -1 

I\ ” -.-- 

6.954 12.7381 

8.692 13.4221 

I 

37.761 

14.8671 

51.895 

120.431) 

.i 

REAR SPAR PLANE 

11,677 

f (4.5971 

~ ““I~“~“Y3r”ILtnb 

IL---- 30.793 (12.1231 

---__--- 43.411 I1 7.091) 

OUTaOARO TWO.SEGMENT FLAP 

‘- LOW-SPEED AILERON 

FIGURE 3. WING (w,,) DIAGRAM 
SH = 0.1298 ITI* 11.397 FT*) 

Ffl = 3.80 

x = 0.350 

SWEEPCH = 30’ 

4 

r = lo.o” 

DIMENSIONS IN CENTIMETERS (INCHES) 

MODEL SCALE 

9.596 ($.778)1 

-Y = 231.87 (91.287) 

0.781) 

PlVflT . . .U, -,.I” AYIC I”.“” KR I\6 

-PERCENT CR 

4.536~J .7K6) 

ABOVE FRP 

YH = 16.11 (6.344) 

\MODEL ACTUAL TRAILING EDGE 

(CUT BACK TO ACHIEVE 0.03 (0.010) 

THICK TRAILING EDGE) 

FIGURE 4. HORIZONTAL STABILIZER (HIA) DIAGRAM 

Sv = 0.09850 III* (1.0663 FT*) 

pi = 1.600 

h = 0.35 

SWEEP Cv = 35O 

4 

DIMENSIONS IN CENTIMETERS (INCHES) 

MODEL SCALE 

t12.87 

(5.065) 

MACV = 26.731 (10.524) 

THEORETICAL 

TRAILING EDGE 

ACTUAL MODEL TRAILING 

EDGE (CUT BACK TO 

ACHIEVE 0.03 (0.010) 

THICK TRAILING EDGE) 

36.759 

(14.472) 

FIGURE 5. VERTtCAL STABILIZER (V,A) DIAGRAM 

19 
Flow-through nacelles (Figure 6) from a DC-10 model were used and were 

attached to the wing by pylons. The pylon plane of symmetry had a 1.8O 

toe-in relative to the airplane plane of symnietry (measured in the FRP) and 

was perpendicular to the FRP with the wing in a rigged position with a 

dihedral angle of 4.05'. Nacelle strakes were attached to the nacelle for 

most tests. 

DIMENSIONS IN CENTIMETERS (INCHES) 

MODEL SCALE 

-I I- 4.30 (1.69) 

I-b- 3.25 PERCENT CHORO 

ENGINE CENTERLINE 

AT+1.6’lNClDENCE 

TO THE FRP 

LEXISTING DC-10 

GE NACELLE (N2A) 

FIGURE 6. NACELLE/PYLON (Na Pm) DIAGRAM 

The nose gear simulated the DC-10 nose gear in structure and location. The 

main landing gear simulates the airplane gear configuration with oleos 

extended. Extended main gear wheel well cavities were not simulated. A 

retracted main landing gear configuration was also provided. 

The definitions of gap, overhang (O.H.), and deflection used to position the 

leading-edge high-lift devices are illustrated in Figure 7. The deflection 

angles were measured in a streamwise plane oriented normal to the wing 

reference plane (WRP). Definitions for main and aft flap gap, O.H., and 

deflections are shown in Figure 8. The same definitions were used for both 

the flaperon and the main flap. The variable test positions tested are 

defined and identified in the grid notations table of Appendix C. 

20 
CLEAN WING 

MAX LENGTi-l LINE 

VCK. FCK 

/ bRACKET 

MLL 

FIGURE 7. LB-486 LEADING EDGE DEVICE GAP, OVERHANG, AND DEFLECTION DEFINITIONS 

FIGURE 8. LB-486 FLAP GAP, OVERHANG, OVERHANG, AND DEFLECTION DEFLECTION DEFINITIONS DEFINITIONS v 
LB-486B,C INSTRUMENTATION 

Aerodynamic forces on the model were measured using the Ames Task Mark II 

10.16-cm (4-in.) diameter internal balance at the Ames l&Foot Pressure Wind 

Tunnel (LB-486B test). For the NASA Langley V/STOL Wind Tunnel 

(LB-486C test), the balance used was the Langley 5.08-cm (Z-in.) diameter 

internal balance. 

In the Ames test, electrolytic alignment bubbles housed in the fuselage nose 

were used to measure the angle of attack of the fuselage reference plane. 

From angles of attack of -6O to O", the model was pitched by the 

external pitch drive. From O" to +lO" angles of attack, the fuselage 

was pitched using the fuselage internal pitch drive while maintaining the 

balance at 0'. For angles of attack of 10' to 34O, the fuselage was 

pitched using the external pitch drive with a loo angle maintained between 

the balance axis and the fuselage axis. 

In the Ames test the horizontal stabilizer incorporated remote drive and 

dual-position potentiometer for changing tail incidence during a run. In 

the NASA V/STOL test, a NASA-furnished electronic inclinometer was used to 

determine angle of attack. The horizontal-tail incidence in the V/STOL test 

was set at O". 

LB-486B,C MODEL INSTALLATION 

The model was installed in the NASA Ames 12-Foot Pressure Wind Tunnel on the 

tandem support system shown in Figure 9. The model was pivoted about the 

main strut pivot point and was powered by the aft pitch strut. The entire 

strut system was nonmetric (i.e., air loads on the strut are not sensed by 

the balance). The struts entered the fuselage as far aft as practical to 

minimize the aerodynamic interference effects on the model. 

22 
BALANCE 

CENTER 

TS 

306.616 

(120.715) 

I 

AMES TASK 

MK II BALANCE 

1 DATAREFERENCE 

CENTER 

oiM~r~s~0r4s IN CENTIMETERS (INCHES) 

MODEL SCALE 

- - -- 

s 

-- 

- 29.57 2.985 

(11.64) (1.175) 

FIGURE 9. MODEL ItiSTALLATlON IN THE NASA AMES 12-FOOT PRESSURE WIND TUNNEL 

The same support system (Figure 10) was utilized during the NASA Langley 

V/STOL test program. It was adapted to the existing V/STOL Tunnel 

structure; extensions for the main and pitch struts were added to the basic 

tandem strut system. The extensions permitted the model to be located near 

the vertical position of the tunnel centerline. 

REVIEW OF PHASE I RESULTS 

During Phase I, the aerodynamic characteristics of the clean wing, VCK, 

slat, and flaps were defined experimentally. The lift and pitching-moment 

curves for the clean wing are shown in Figure 11. These curves indicate 

that the cruise wing, as defined for Phase I, was subject to outboard stall, 

although it is likely that the curves overstate the tendency for stall 

because of the Reynolds number effect. Because of the short tip chord of 

the wind tunnel model, the highest Reynolds number condition resulted in a 

tip chord Reynolds number of only 1.9 million. Figure 12 shows that higher 

23 
TS = 506.43 (200.17) 

MS = 160.28 (63.102) 

MS = 65.301 (25.709) I 

LANGLEY 748 BALANCE 

BALANCE CENTER- /m I 

8AYONET- 

,i , ~P~TCHSTRUT 

MAIN STRUT 

EXTENSION 

-I- 

($ MAST SUPPORT 

M 

I 

I 

PITCH STRUT 

c= EXTENSION 

I 

TOP SURFACE 

V/STOL TUNNEL FLOORPITCH DRIVE 

MECHANISM 

I DIMENSIONS IN CENTIMETERS (INCHES) 

MODEL SCALE 

FIGURE 10. MODEL INSTALLATION IN THE NASA LANGLEY V/STOL WIND TUNNEL 

stall angles and larger values of section C LMAX 

's for the outboard wing 

panel might have been obtained if the test could have been made at a higher 

Reynolds number. Later high-aspect-ratio supercritical wing designs have 

shown improvements in stall angles and C LMAX' 

24 
l . 

MODEL LB466A i Ii SYM RUN 

f3 

CONFIGURATION S-N-- P-m te _ V. _ H. _ 

” 85 a 1 OFF t 0 1 23 i 1 1 ZA ZA 1A 1A 1A ;i 0. YOO p& 

MACH = 020 

3. 50- RNMAC = 5.12 x lo6 

I 

3. oo-; 

2. 00 

I. 50 

I. 00 

0. 50 

0 

0” 

- 

-5; 

0” B 

8 -0. 50 / 

I I I 

-10 25 30 

I- 

-0.200 

-0.300 

-0. L100 

I I I I I I 

5 IO I5 20 25 30 

RN6LE OF RTTRCK-DEG 

I3 

ANELE OF RTTRCK-DEG 

.~- 

LIFT AND PITCHING MOMENT 

FIGURE 11. TAIL-ON CHARACTERISTICS FOR THE CRUISE WING WITH NACELLES, PYLONS, AND STRAKES ATTACHED 
MODEL LB-488A 

CONFIGURATION S, NzA PPAZIA 

MACH = 0.20 

1. 

1 .c 

FIGURE 

1 

12. EFFECT OF 

2 

REYNOLDS 

5 10 

RN 6 

MAC x 10 

NUMBER ON CLEAN-WING 

20 

SECTION MAXIMUM 

50 

LIFT 

26 
Figures 13 and 14 show the lift and pitching-moment characteristics for the 

primary VCK and slat configurations tested. While the C 

LMAX 

and L/D ratio 

for the slat configurations were marginally better than those of the VCK 

configurations, use of the VCK resulted in superior stall characteristics. 

Configurations including slats exhibited both pre-stall and post-stall 

nose-up tendencies. While the VCK configurations showed post-stall nose-up 

trends, the pre-stall characteristics were good. Nearly all of the work 

accomplished on this model during Phase II was directed toward improving the 

low-speed stall characteristics by making adjustments in leading-edge device 

position and type. 

The trailing-edge flap studies of Phase I indicated that the changes in 

performance due to gap and overhang variations were not as significant as 

the corresponding variations for the leading-edge devices. As expected, the 

two-segment flap was superior to the single-segment flap in C LMAX and flap 

lift increments. Trimmed polar comparisons indicated that the 

single-segment and two-segment flaps resulted in equivalent L/D envelopes 

for takeoff flap settings. For equivalent values of approach speed, the L/D 

values for the two-segment flap were superior to those of the single-segment 

flap. Because of these definitive results, little additional flap 

optimization work was conducted on the wide-body model during Phase II. In 

addition to the high-lift work, Phase I testing also defined the 

effectiveness of the spoilers and ailerons. 

27 
3. 50. 

3. 00 

2. 50 

2. 00 

I. 50 

I. 00 

C&O 

MODEL LB466A 

CONFIGURATION S4HIAVIA 

MACH = 0.20 

RNI\IIAC = 5.12 x lo6 

6 VCK = 45El45G 

‘FLAP = 5CllOB 

a 

B 

00 

0 

; 1’0 1’5 ;0- 

- 

25 

RNGLE OF RTTRCK-DEG 

I Qc@3 

-5 

-0. 100 

-0. 200 

0 o Q 

-0. 300 

-0. qoc 

-0.5oc 

-0. hoc 

-0. 7oc 

i H 

OFF 

O0 

-5O 

RUN 

116 

121 

122 

0 

I I I 1 I I 

5 

RNGLQO:‘flTTRCK-D;; 

20 25 30 

LIFT AND PITCHING MOMENT 

FIGURE 13. TAIL-ON AND TAIL-OFF AERODYNAMIC CHARACTERISTICS OF THE VCK WITH TWO-SEGMENT TAKEOFF FLAPS CONFIGURATION 
3.50-, 

3. 00 i 

I 

2.50-l 

MODEL LB-466A 

BASIC CONFIGURATION S4HIAVIA 

MACH - 0.20 

RN MAC = 5.12 x 10’ 

6 - 15D125D 5 

SLAT Y 

6 = 5CllOB 2 

FLAP I: 0.200- 

jo 

:J;r -o.boo 

-0. 700 i 

-._.-- 

LIFT AND PITCHING MOMENT 

c: 

FIGURE 14. TAIL-ON AND TAIL-OFF AERODYNAMIC CHARACTERISTICS OF THE SLAT WITH TWO-SEGMENT TAKEOFF FLAPS CONFlGURAtltJN 
LB-486B,C RESULTS AND DISCUSSIONS 

Most of the work on the wide body model during Phase II was directed toward 

improving the pitching-moment characteristics of the wing, without causing 

an excessive loss in C 

LMAX' 

The approach consisted of either increasing 

the stalling angle of the outboard wing panel, or tuning the stall angle of 

the inboard wing to be just below that of the outboard wing. Additionally, 

to prevent post-stall pitch-up, it was desirable that the stall inboard be 

due to separation at the leading edge of the high-lift device, thereby 

increasing the rate of lift loss inboard relative to that outboard. 

Configurations tested included a VCK with a reduced deflection, trimmed 

slats inboard, a normal-chord and a short-chord FCK, a differential flap 

deflection, and a two-segment flaperon. In addition to the study of these 

configurations designed to improve C 

LMAX 

and/or pitching-moment trends, 

the improvement in takeoff L/D performance due to sealed slats was 

evaluated, the penalty associated with use of a high-speed aileron was 

determined, and data obtained at the Langley and Ames tunnels were compared. 

Reduced VCK Deflection 

Phase I results (LB-486A) showed equivalent C 

LMAX 

values for the slat and 

VCK configurations. However, the lower minimum pressure coefficients on the 

VCK indicated that a reduction in deflection might delay leading-edge 

separation and result in increased maximum lift. A VCK deflection of 

&VCK = 33' compared to the Phase I value of 6VCK = 45" was 

therefore selected for the LB-486C test at the NASA Langley V/STOL 

Facility. Results of this test indicated that it was not possible to obtain 

increased C 

LMAX 

due to the low Reynolds number (1.14 million) available in 

this tunnel. Further examination of the configuration was made at a higher 

Reynolds number (5.89 million) during the Ames 12-Foot Tunnel entry 

(LB-486B). The same results as in LB-486C were observed. The reduced 

deflection resulted in a lower outboard stall angle than the 45O 

deflection. The basic 45', 33', and 45O/33O (inboard/outboard) VCK 

deflection lift and pitching-moment data are shown in Figure 15. The 

corresponding drag values indicated L/D values at 1.3Vs of 9.52, 10.0, and 

9.0 for the 45', 33', and 45O/33O VCK deflections, respectively. 

30 
3. 50. 

3. 00. 

2. 50. 

2. 00. 

I. 50. 

8 

I. 00. 

0. 50. 

I 

-5 * 

3 

-0. 50. 

NSTC 

MODEL LB4866 ;YM 1 RUN t 

CONFIGURATION B 2A w3B ‘2B N2A ‘2A Z 1A 

MACH = 0.20 

RN MAC = 5.12x lo6 

6 FLAP = 25Kl12C 

B 

(P 

P 

I I I I I 

5 IO 15 20 25 

RNGLE OF RTTRCK-DEE 

q 

-0.200 

-I 

-0. 300 

-0. YOO 1 

RNGLE OF RTTRCK-DEG 0 0 

r3 

A. LIFT AND PITCHING MOMENT 

FIGURE 15. EFFECT OF VCK DEFLECTION 
C. 

9. 0 

8. 0 

7. 0 

x 

. 

i 

: 

$ b.0 

L 

b 

: 

3 5.0 

.Y 

4. c 

3. c 

2. c 

I. c 

--ed 

04 t 

-9 

III- 

,- 

I- 

,- 

,- 

b-f3. ( 

I I I I I ! I I I I 1 I I 1 I I 1 

10 0. 04 0. ox 0. I.2 0. lb 0. 20 0. 24 0. 2x 0. 32 0. 3b 0. 40 0. 44 0. 4s 0. 52 0. 5b 0. b0 ‘0. b4 0. bE 

NSTC DRRG COEFFICIENT 

0. 72 1 

0. DRAG 

FIGURE 15. EFFECT OF VCK DEFLECTION 
Sealed Slats 

In the Phase I LB-486A tests, a landing slats/takeoff flaps combination was 

investigated since it would simplify the high-lift system mechanically to 

have only one slat position for both takeoff and landing. The results 

showed, however, that the landing slat reduced L/D when used with either a 

clean trailing edge (GFLAp = 0') or the basic takeoff flap deflection 

(&FLAP = 5O/lOO). To improve the L/D for this combination a sealed 

(i.e., zero gap) inboard and outboard slat configuration was investigated. 

The configuration was tested first with a 5O slat deflection inboard and a 

20' deflection outboard. Then because previous analysis had shown a 

retracted slat might improve the pitching-moment characteristics, it was 

also tested with a 0' deflection inboard and a ZOO deflection outboard. 

The results are presented in Figure 16. Because the loads on the sealed 

slat were expected to be high, it was not tested at the high Reynolds 

number. The results indicate that, as expected, the 50/20° 

configuration had adverse pitch-moment characteristics. These were improved 

by retracting the inboard slat, without reducing C 

LMAX' 

Also shown in Figure 16 is the landing slat configuraton with takeoff 

flaps. The CLMAX penalty associated with the sealed slat is obvious. 

Figure 16 shows the O"/200 slat configuration gave slightly higher L/D 

than the 50/20° slat configuration, tail-on. Tail-off L/D's for clean, 

sealed, and slotted configurations are compared in Figure 17. The improved 

tail-off L/D values for the sealed configuration at 50/10° flap 

deflection are illustrated. High Reynolds number data for the clean 

trailing edge with sealed slat configuration were not obtained. 

An inboard sealed slat deflection of 5' was tested with landing flaps and 

an outboard landing slat position. Results indicated a substantial C 

LMAX 

degradation and post-stall nose-down pitching-moment trends (Figure 18). 

LB-486A testing included a 15' inboard sealed slat position; the results 

showed no adverse effects on C 

LMAX 

and no change in pitching-moment 

characteristics. An inboard sealed or small-gap slat configuration at an 

intermediate inboard slat deflection is a candidate for future low-speed 

studies. 

33 
“. I..” 

MODEL LB486B I6 RN 6 SYM RUN 

CONFIGURATION I3 2A w3EI NPA ‘ZA’IA H1A “IA =I 

MAC SLAT 

H 0. 300- 

MACH = 0.20 k w 2.89 x lo6 5 SEALED120 SEALED 0 42 

3. 50 s 

1 6 = 5CllO 5 82.89 511x 

lo6 cl 

FLAP 

F 0.200- lo6 CLEAN/SO 15 SLOTTED125 SEALED SLOTTED V 461 

V 

-0. $1 00 0 

B 

i 

0 RNGLE OF 

-0. 200 @cl 

RTTRCK-DEE 0 

0 

0 

vv 

00 

3 

v 

v 

0 

v 

-0. 500 1 0 

q 00 

q 

-0. boo i I7 

1 

I3 

q 

-0. 700 

RNGLE OF HTTRCK-DEG 

-0. 50 

-0. 800 J 

NSTC 

v 

A. LIFT AND PITCHING MOMENT 

FIGURE 16. EFFECT OF SEALED SLATS 
9. o- MODEL LB-‘tXb B 

RN MAC 6 SLAT SYM RUN 

8. o- . 

2.89 x lo6 5 SEALED/PO SEALED 0 42 

2.89 x lo6 CLEAN/PO SEALED Cl 46 

7. o- ; 5.11 x lo6 15 SLOTTED125 SLOTTED V 41 

4. 0 

3. 0 

2. 0 

P 

q @ 

q 0 

$3 

0 

w 

El 

0 

v v v 

V V 

v 

v 

v 

0 

0 

0 

q q 

0 0 

q 

0 0 

El q Q 

I I I I I 1 I , , I 1 1 

0. 2Lt 0. 28 0. 32 0. 3b 0. 40 0. 44 0. 4x 0. 52 0. 5b 0. b0 : 0. b’t 0. b8 ( 

NSTI: DRRE COEFFICIENT 

1. ; 

I 

72 

B. DRAG 

FIGURE 16. EFFECT OF SEALED SLATS 
l8- 

lb12- 

e 

J io8- 

bEl 

B 

q 

MODEL LB486 

CONFIGURATION B 2A w3 ‘2B N2A ‘2A ‘IA 

MACH = 0.20 

RN MAC = 2.58 x lo6 

v 

&? 

0 

El 

El 

v q V 

-B v 

q v 

0 0 a 

El 

q 

v 

0 v 

v 

F 

v 

v 

v 

v 

TEST 6 SLAT 6 FLAP SYM RUN 

LB-486A CLEAN/CLEAN 010 D 24 

LB-486A 15 SLOTTED/25 SLOTTED 5/10 0 178 

LB-486B CLEAN120 SEALED 5110 v 51 

0 

v 0 

v 0 

v 

v 

v 

0 

0 

0 

0 

0 

0 

0 

0 

0 

0 

0 

0 

I I I I I I I 1 I I 

‘%O 0.2 O.L( 0.b 0.8 I.0 1.2 I.L1 I.b I.8 2.0 2:2 

I I I I I I 

2 2. L1 2. b 2. 8 3. 0 3. 2 3. q 

-21 

LIFT COEFFICIENT 

b 

FIGURE 17. L/D COMPARISONS FOR CLEAN WING, SEALED SLAT, AND LANDING SLAT CONFIGURATIONS 
-0, 50 

NSTC 

MODEL LBQB6B 

CONFlGURATliIN 6 

5 

2A w3B ‘2B N2A ‘2A ‘IA H,A “IA 5 6 

:: 

SLAT SYM RUN 

0.300- 

MACH = 0.20 :: 

RNhlAC = 1.14x lo6 ki 15D125D 0 23 ” 

6 = 25112 2 0.200- 5 SEALEDl25D 0 24 

FLAP 2 

:” 

I: 

0000 J 0. IOO0 E 

0 QOOO =I 

a 

io 1’5 i0 25 j0 

RNGLE OF RTTACK-DEG 

El 8 

q -0.200- 

-0. boo- 

, I I I I 

5 IO I5 20 25 

-0. 700- 

RNGLE OF RTTRCK-DEG 

-0.800 I I3 El 

A. LIFT AND PITCHING MOMENT 

FIGURE 18. EFFECT OF INBOARD SEALED SLAT WITH LANDING FLAPS 
8. 0. 

7. 0 

x 

* 

; 

: 

I: b. 0, 

L 

5 

3 

5 25.0 

0 MODEL LB-486B 0 0 0 

0 

0 

0 0 

0 

0 B 

q 

B 

El 

El 

El 

NSTC DRRG COEFFICIENT 

B. DRAG 

FIGURE 18. EFFECT OF INBOARD SEALED SLAT WITH LANDING FLAPS 
Fixed-Camber Krueger 

A fixed-camber Krueger (FCK) is an attractive high-lift device option, 

especially inboard, because of its mechanical simplicity and the need to 

stall the inboard wing panel just before the outboard panel stalls. The 

capability of a very efficient slat or VCK is not needed. As shown in 

Figure 19, the full-span FCK produced lift and pitching-moment 

characteristics equivalent to those of the full-span slat and full-span VCK 

configuration. Use of an FCK inboard with a slat outboard, however, 

resulted in improved pitch characteristics (Figure 20). Even though the 

FCK/slat combination caused pitch-up to start at a lower angle of attack 

than the FCK/FCK combination, pre-stall nose-up tendencies were greatly 

reduced, and could possibly be eliminated with additional tuning. 

Post-stall characteristics continued to be unsatisatifactory, indicating a 

lack of leading-edge separation on the FCK. 

To further improve pitching-moment characteristics, a short-chord FCK was 

fabricated and tested during the LB-486B series. The chord ratio for this 

device was 0.068, extrapolated to the side of the fuselage, and 0.105 at the 

leading-edge break (pylon position). The comparable values for the slat 

were 0.1803 and 0.1295, respectively. The bulb shape was tailored such that 

an inboard, leading-edge stall would be obtained. FCK deflections of 50' 

and 70' were evaluated with zero gap and overhang. Examination of the 

trailing-edge pressures indicated that a premature inboard stall was being 

obtained. Favorable pitch characteristics at stall were obtained 

(Figure 211, but at the expense of a substantial reduction in C 

LMAX 

values 

of -0.457 and -0.412, respectively, for the two FCK deflections. Shims were 

fabricated at the tunnel to obtain a small gap and negative overhang for 

this leading-edge device. The best FCK/slat configuration resulted in 

higher maximum lift values and better pitching-moment trends then did the 

full-span slat configuration (Figure 22). Tail-off drag values indicated 

L/II values at 1.3Vs of 9.71 and 9.77 for the FCK and basic slat 

configuration, respectively. 

39 
MODEL LB-499C 

CONFIGURATION B 2A w3Fl ‘ZB N2A ‘2A ‘IA 

MACH = 0.20 

RN MAC = 1.14x lo6 

6 FLAP = 25112 

-0 I 1 1 1 I I I 

4 0 Lf 8 12 16 20 2L1 28 

1 ANGLE OF ATTACK (DEG) 

as It 

.3 

.2 

l- 1 

-0 

ANGLE OF ATTACK (DEG) 

-. 1 

A. LIFT AND PITCHING MOMENT 

FIGURE 19. EFFECT OF LEADING EDGE DEVICE WITH LANDING FLAPS 
I P 

18- 

16 - 

111- 

12 - 

MODEL LB-486C 

10 - 

LID 

8- 

8- 

2 

t 

I 

OO 

I I I I I I I I I I I I I t 

.2 .It .8 -8 1.0 1.2 1.q 1.6 1.8 2.,-J 2.2 24 2.8 2.8 310 

LIFT COEFFICIENT 

B. LIFT-DRAG RATIOS 

FIGURE 19. EFFECT OF LEADING EDGE DEVICE WITH LANDING FLAPS 
R 

3.2 

2.8 

MODEL LB-466C 

CONFIGURATION B 2A w3B ‘2B NZA ‘2A ‘,A ‘IA H1A 

MACH = 0.20 m 

= 1.14 x lo6 

.> 

RN MAC 

6 FLAP = 25112 

.2 

I 

-0 I I I I I I I 

-.q 0 Lt 8 12 16 20 24 28 

1 ANGLE OF ATTACK-DEG -.6 

-.Lf 

-.7 

I I I I I I I 

q 6 12 16 20 2Y i8 

ANGLE OF ATTACK - DEG 

A. LIFT AND PITCHING MOMENT 

FIGURE 20. COMPARISON OF FULL-SPAN FCK AND FCK/SLAT COMBINATION 
MODEL LB486C 

18 - 

16- 

14 - 

12 - 

L/D 

10 - 

2- 

OO 

I , I I I I I I I I I I I I 

.2 .q .6 

I 

-8 1-o 1-2 I.'! 1.6 1.6 2.0 2.2 2.q 2.6 2.8 3.0 

LIFT COEFFICIENT 

B. LIFT-DRAG RATIO 

FIGURE 20. COMPARISON OF FULL-SPAN FCK AND FCK/SLAT COMBINATION 
3. 50 

3. 00 

2. 50 

2. 00 

I. 50 

0 

I. 00 

0. 50 

I 88 

MODEL LB-466B 5 

CONFIGURATION B 2A w36 ‘PI3 N2A ‘2A ‘IA ‘lAHIA 

:: 

Ii 0.300 

MACH = 0.20 k 

RN = 6.1 x IO6 :: 

MAC u 

8 

81 

-0. 200 

1 

-0. i oo0 

-0. +00-Q Q 

01 

@Q 

-0. 500- OR 

8 

I I 1 1 I I 

-5 5 IO 15 20 25 

-0. 700 

i 

J RNELE OF RTTRCK-DEE 

-0. 50 

-0.800 1 

NSTC 

RNGLE OF RTTRCK-DEG 

A. LIFT AND PITCHING 

FIGURE 21. GRID STUDY FOR SHORT-CHORD FCK/SLAT COMBINATION 
7. o- 0 

MODEL LBQBBB 

x. o-’ 

e 

Y. oI 

I 

* 0.00 

I I 1 I 1 I I I I I 1 I I I I I I 

OLt 0. OS 0. ox 0. 12 0. lb 0. 20 0. 24 0. 28 0. 32 0. 3b 0. YO 0. YY 0. $X 0. 52 0. 5b 0. b0 0. bY 0. bZ 0 

NSTC DRRG COEFFICIENT 

B. DRAG 

FIGURE 21. GRID STUDY FOR SHORT-CHORD FCK/SLAT COMBINATION 

72 
MODEL LB-4BBB 

v-66 

CONFIGURATION B 2A w3B ‘2s N2A ‘2A ‘IA “IA HIA t 

MACH = 0.20 :: 

=I 0. 300 

RN MAC 

= 5.11 x lo6 k 

8 

3. 50 _ 

3. oo2.50- 

a 

NSTC 

RNGLE OF P.TTRCK-DEG 

-0. 5c 

0 

0 -0. 20( 

-0. 3oc 

; 

7- 

IO ;5 20 25 30 

RNGLE OF RTTRCK-DEG 

0 0 

0 

A. LIFT AND PITCHING MOMENT 

FIGURE 22. COMPARISON OF FULL-SPAN SLAT AND SHORT-CHORD FCK/SLAT COMBINATION 
MODEL LB486B 

6.0 GI 

q 

FCKISLAT 70Di26D 0 35 

3.0 - 

2.0 - 

1 .o 

I. 

-0.e 

NSTC 

- 

1 , 1 1 --- 1 \ I 

-. 

I 1 I 1 1 , L 

0.04 0.08 0.12 0.16 0.20 0.24 0.28 0.32 0.36 0.40 0.44 0.48 0.52 0.56 0.60 0.64 0.68 

DRAG COEFFICIENT 

B. DRAG 

FIGURE 22. COMPARISON OF FULL-SPAN SLAT AND SHORT-CHORD FCK/SLAT COMBINATION 
Slat Trim Effects 

The lift and pitching-moment characteristics for the revised slat trim are 

presented in Figure 23. The basic trim consisted of a side-of-fuselage 

inboard trim and a sealed over-the-pylon configuration (i.e., continuous 

,over the pylon). This base case resulted in a C 

LMAX 

value of 3.2. 

Figure 23 also illustrates two other trim variations which showed a C 

LMAX 

reduction of approximately 0.20. For the first variation, the slat trim was 

moved outboard 2.25 cm (1 in.) from the fuselage side. This resulted in 

improved pitch characteristics at the stall angle, but pitch-up at 

post-stall conditions. In the second variation, in addition to the revised 

inboard slat trim an over-the-pylon island (i.e., undeflected slat) trim was 

tested. Pitching-moment characteristics similar to those of the basic trim 

resulted but with reduced magnitude of pitch-up. Small effects were noted 

on L/D performance for the two slat-trim revisions. Examination of 

Figures 22 and 23 indicates a lower C 

LMAX 

and more adverse post-stall 

behavior for the slat trim configuration than the short-chord FCK. 

High-Speed Aileron 

In order to determine the benefit of a flaperon, a configuration using a 

high-speed aileron in place of the flaperon was tested at the maximum 

landing flap deflection of 35O/12'. The results indicated a reduction 

of 0.315 in CLa = o and 0.216 in C 

LMAX* 

The drag increase at 1.3Vs 

was 0.008. High-angle-of-attack pitch characteristics were essentially 

similar to those of the basic configuration. 

Two-Segment Flaperon Replacement 

For several runs, the single-segment flaperon was replaced with a 

two-segment flaperon. The effects of the change were evaluated at landing 

and takeoff flap deflections. The increases in corresponding C 

were 0.061 and 0.039, respectively. LMAX 

values 

Small changes in pitching moment were 

also indicated. The drag values indicated essentially no change due to the 

two-segment replacement for the single-slot flaperon. 

48 
I 

3. 50- 

3. ooMODEL LB-4666 

CONFlGURATlOti B 2A w313 ‘2B N2A ‘2A ‘IA HI A “IA 

MACH = 0.20 

RNMAC = 6.11 x lo6 

6 FLAP = 26112 

8 

c 

z 

=1 l-4 0.300- 

: 

i?i 

2 0. 200- 

r;! 

? 

x 0. IOOz 

b 

/SIDE OF FUSELAGE TRIM (BASIC) 

OUTBOARD TRIM 

CONTINUOUS OVER 

PYLON (BASIC) 

0 

2 

2. ooz 

w I 

=1 I 

e % 

ki 8-l 1.50- 8 

t 

t: 

BASIC (SIDE OF FUS AND OVER PYLON) 

I . 1 I I I 1 

-5 cr 5 IO I5 20 25 

RNELE OF RTTRCK-DEG 

-0.50- 

NSTC 

RNGLE OF RTTRCK-DEG 

-0. 500- oooooo 

0 ITI 

oooooD~ 

-0. bOO- 

-0. 700 

I 

-0. zoo I 

0 

El 

0 

-- -- 

A. LIFT AND PITCHING MOMENT 

FIGURE 23. EFFECT OF SLAT TRIM WITH LANDING FLAPS 
9. 0, 

x. 0 

Y. 0 

3. 0 

2. 0 

I. 0 

SYM RUN SLAT TRIM CONFIGURATION 

@ 

I 

09 . 0.00 

I I I I I I I I I I 1 I I I I I I 

0. OLt 0. OX 0. 12 0. lb 0. 20 0. 2Y 0. 2x 0. 32 0. 3b 0. Lto 0. w 0. YX 0. 52 0. 5b 0. ho 0. bY 0. bX c 

NSTC DRRG COEFFICIENT 

72 

B. DRAG 

FIGURE 23. EFFECT OF SLAT TRIM WITH LANDING FLAPS 
Differential Flap Deflection 

A 35'/12' (main flap/auxiliary flap) inboard flap deflection combined 

with a 25'/12O outboard flap deflection was also tested to determine the 

effect on the low-speed characteristics. Results compared with those of the 

basic 25O/12' two-segment flap deflection indicated a small reduction in 

C LMAX (-0.046) and slightly more positive pitching moments. The increased 

inboard flap deflection did not produce a smaller inboard stall angle and 

the associated stall improvements. The differential flap deflection did 

result in a drag increase of 0.0180 for the C, range of interest. 

Ames 12-Foot and Langley V/STOL Tunnel Comparisons 

During the Phase I wind-tunnel tests in the Ames l2-Foot Pressure Tunnel, 

several configurations were tested at high Reynolds number as well as at 

atmospheric conditions. Two of these configurations were also tested in the 

Langley V/STOL facility for comparison. The tandem strut support system was 

utilized in both cases. Figure 24 presents the lift and pitching-moment 

comparison at the atmospheric condition for the slat with two-segment 

takeoff flap configuration. The data presented have been corrected for 

tunnel wall effects, but not for strut tare effects since these would be the 

same for both wind tunnels. Good agreement between the Ames and Langley 

data is shown for the lift coefficient up to the angle of attack for stall. 

Sane differences are noted in the post-stall region. The pitching-moment 

data show differences for most of the angle-of-attack range. This was also 

typical of the VCK configuration used for comparison. Comparison of the 

drag characteristics indicated differences of 0.0050 to 0.0070 for the 

configurations evaluated. The Ames wall corrections are considered a 

possible source of these differences. 

51 
MODEL LB4B6B & C 

3.2 1 

MACH = 0.20 

RN MAC = 1.14x lo6 

HORIZONTAL TAIL-OFF 

LANDING GEAR OFF 

2.8 

2.4 

, * 

1.2 

0.8 

.4. p6 

-FY- d 4 6 li 16 io i4 & 

I 

RNGLE OF RTTRCK-DEG 

-0.4 

RNGLE OF RTT 

-0.5 - 

-0.6 - 

SLAT (15’/25’) +TWO-SEGMENT FLAP (5°/100) 

LIFT AND PITCHING MOMENT 

FIGURE 24. AMES 12-FOOT AND LANGLEY V/STOL COMPARISON 
PART II 

NARROW BODY 

ATMR-TYPE MODEL 

L-B-507A.M’ObEL 

INSTALLED IN AMES 

i2-FOOT PRESSURE TUNNEL 

53 

PART II 

NARROW BODY ATMR-TYPE MODEL 

LB 507A MODEL DESCRIPTION 

i A 5.59-percent-scale full-span model of the ATMR aircraft was used for this 

program. This model is shown in Figure 25. The configuration notation 

data, dimensional data, and grid position definition are presented in 

Appendixes D, E, and F, respectively. The model included a 

high-aspect-ratio supercritical wing, variable-position leading-edge slats, 

an inboard short-chord FCK, two-segment trailing-edge flaps, wing and 

high-lift surface pressure instrumentation, and a remotely driven horizontal 

stabilizer. The outboard ailerons and wing spoilers also had deflection 

capabilities. The model instrumentation was equipped with the Douglas 

internal pitch system. This system was used in conjunction with the Douglas 

tandem support system and the Task MK IIC internal strain-gage balance. 

The model fuselage utilized the LB-506A (high-speed EET model) nose section 

and glass fiber wing/body fillet. These parts were combined with a new 

aluminum centerbody and aft section. The constant-diameter hollow center 

section was machined on the upper and lower surfaces and internally to 

provide clearance for the Douglas 10.16-cm (4-in.) balance housing and 

internal pitch system. Other instrumentation housed in the fuselage 

included two 6-pat scanivalve modules in the nose, two electrolytic bubbles 

measuring the angle of the balance axis, and an electrolytic bubble pack to 

measure the fuselage'angle of attack. 

The wing for this model (Figure 26) consisted of right- and left-hand panels 

which were joined together and to the fuselage by means of a wing splice 

plate. The wing had removable leading and trailing edges to allow for the 

attachment of high-lift devices, and had movable control surfaces. The wing 

also included pressure instrumentation at four spanwise locations, and had a 

trailing-edge pressure port at one inboard span location. A diagram of the 

high-lift system and the lateral control surfaces is provided in Figure 27. 

55 
MODEL LB-507A 

DIMENSIONS IN CENTIMETERS (INCHES) 

MODEL SCALE UNLESS OTHERWISE SPECIFIED 

X MAC = 

TRAPEZOIDAL AREAS: 

S REF = 0.464 III’ (5.000 FT*) 

sli = 0.113 rn* (1.231 FT*) 

% = 0.086 III* (0.931 FT*) I 

tY- = 136.976 

c/4 (53.888) - 

-L = 100.952 - 

V (39.745) 

I 

13.485 

(5.309) 

FIGURE 25. LB-507A MODEL THREE VIEW 
l 

12.670 

(4.988) 

L 

MODEL LB-507A 

TRAPEZOIDAL WING CHARACTERISTICS 

S REF = 0.464 m* (5.000 FT*) 

A c/4 = 26.OOODEG 

T.R. = h = 0.275 

#I = 11.10 

MAC = 2.103 cm (8.922 IN.) 

b/2 = 113.532 cm (44.698 IN.) 

DIHEDRAL = I- = 5.000DEG 

DIMENSIONS IN CENTIMETERS (INCHES) 

MODEL SCALE UNLESS OTHERWISE NOTED 

FIGURE 26. LB-507A WING IW,,) 
1 L INBDSLAT \ UI!l 

MODEL LB-507A 

DIMENSIONS IN CENTIMETERS (INCHES) 

MODEL SCALE 

3 L INBD SPOILER 

X f 1A 

r OUTBD SLAT L,, 

OUTED SPOILER fpA 

FIGURE 27. HIGH-LIFT AND LATERAL CONTROL SURFACES 
'The model utilized the XIB fillet which was developed for the high-speed 

Model LB-506A. The glass fiber fillet was modified on the lower surface to 

provide access holes for the Douglas tandem support system. 

The model was equipped with one set of inboard and one set of outboard 

leading-edge slats. The slats were attached by rigged brackets to a WUSS 

leading edge which was interchangeable with the cruise leading edge. 

Brackets were available to rig the inboard slats at three different 

positions. At one of these three positions, a set of shims could be 

installed between the slat brackets and the wing to provide a fourth slat 

grid position. The definitions of slat gap and overhang are shown in 

Figure 28 (which is Figure 7 repeated for convenience), the various slat 

SLAT 

OVERHANG 

(-I SHOWN 

CLEAN WING 

MAX-LENGTH 

LINE 

FCK 

/ I, BRACKET 

MLL 

FIGURE 28. LEADING EDGE DEVICE GAP, OVERHANG, AND DEFLECTION DEFINITIONS 

59 
deflections and grid positions are provided in Appendix F. The slats also 

contained pressure instrumentation at four spanwise locations. The inboard 

leading-edge slat could be replaced with a short-chord fixed-camber 

Krueger. This FCK could be positioned at two deflection angles with two 

grid positons at each angle. The FCK did not contain pressure 

instrumentation. 

The trailing-edge high-lift system consisted of 80-percent span two-segment 

flaps. The flaps were continuous, with no inboard aileron or exhaust gate. 

They were installed in the desired positions using fixed brackets which 

attached the main flap to the wing and the auxiliary flap to the main flap. 

Each forward flap segment could he installed at four deflection angles, and 

each aft flap segment could be installed at two deflection angles. The 

bracket attachments were such that the aft flap angles were independent of 

the forward flap angles, allowing either aft deflection and grid position to 

be used with all four main flap settings. The exact flap deflections and 

grid positions are given in Appendix F. The cruise configuration model 

utilized the same flap linkage fairings as the cruise wing of the high-speed 

LB-506A. For the flap-deflected case, a new set of fairings was used. The 

new fairing were set in one position relative to the main flap, and 

represented the fairing position for maximum fairing deflection. The 

definitions of the flap gap and overhang are presented in Figure 29. 

+ OVERHAN 

FIGURE 29. FLAP GAP, OVERHANG, AND DEFLECTION DEFINITIONS 

60 
The outboard ailerons on this model, attached with fixed brackets, could be 

manually positioned at several deflection angles. The model was equipped 

with inboard and outboard spoilers, as shown on the control surface diagram 

of Figure 27. On the model, a one-piece bent-plate-type spoiler was used to 

represent the airplane's three inboard panels, and a one piece 

bent-plate-type spoiler was used to represent the outboard three panels. 

A set of landing gear, which included two wing-mounted gear and one nose 

gear, could be installed on the model for use in the landing or takeoff 

configuration. The airplane gear wells and gear doors were simulated on the 

model, and gear well fillers were provided for the gear-up case. 

The horizontal and vertical stabilizers from the high-speed LB-506A model 

were used on this model. The horizontal stabilizer was adapted to a 

remote-drive and position-indication system, and was modified slightly to 

match the new aft fuselage lines. The vertical fin was installed on this 

model such that the exposed area was the same as on model LB-506A. This 

placed the top of the vertical stabilizer at a different height due to the 

change in aft fuselage lines. The dorsal fin was also used; however, the 

contour of the dorsal was changed as shown in Figure 30. Horizontal and 

vertical stabilizer diagrams are presented in Figures 31 and 32, 

respectively. 

Two wing-mounted nacelles and pylons were used on this model. These parts 

were the nacelle/pylon combination previously tested on model LB406A. The 

flow-through nacelle represented that of the Pratt & Whitney Aircraft JTlOD 

engine. The flap-linkage fairing incorporated into the pylon was modified 

to allow the fairing to deflect with the flap. 

LB-507A INSTRUMENTATION 

The instrumentation associated with this model included a six-component 

internal balance, wing static pressure orifices, a remotely driven 

horizontal stabilizer, and an internal fuselage pitch system. The internal 

pitch system and remotely driven horizontal stabilizer required the standard 

Douglas power supplies, control console, and position readout systems. The 

control console also included Douglas bubble-pack monitoring equipment. 

61 
MODEL LB-507A 

D 2A DORSAL L.E. 

7 

FUS (B,,) 

FIGURE 30. DORSAL FIN (DzA) 
sH = 0.114 n? (1.2312 FT’) 

AR = 4.10 

h = 0.350 

SWEEPC, = 30’ 

4 

rH = lo.o” 

MODEL LB-507A 

THEORETICAL 

TRAILING EDGE 

L MODEL ACTUAL TRAILING EDGE 

(CUT BACK TO ACHIEVE 0.0254 cm 

(0.01 INCH) THICK TRAILING EDGE) 

I------- 

:9” 

1c 

HORIZ STAB. 

ORIGIN. 

Y = 229.022 

(90.166) 

- PIVOT AXIS 65.42% C 

Y = 245.209 (96.539s 

2 = 6.759 (2.661) 

DIMENSIONS IN CENTIMETERS (INCHES) 

MODEL SCALE 

FIGURE 31. HORIZONTAL STABILIZER (H ,D) 
% = 0.086 rn2 (0.931 FT2) 

AR = 1.600 

h = 0.35 

SWEEP C, = 35’ 

4 

217.998 

(85.326) Y 

+z” 

t 

MODEL LB-507A 

I 

12.057 

- (4.7469) 

-i 

= 19.829 (7.807) 

= 15.617 (6.1483) 

37.203 

(14.647) 

THEORETICAL 

ACTUAL MODEL TRAILING EDGE 

:“U^T’-K~K~+ZE 0.0254 fo.01 I THICK) I 

VERT STAB. _/ 

ORIGIN 

DIMENSIONS IN CENTIMETERS (INCHES) 

MODEL SCALE UNLESS OTHERWISE SPECIFIED 

FIGURE 32. VERTICAL STABILIZER (V,D) 
Aerodynamic forces on the model were measured using the Ames Task Mark IIC, 

10.16-cm (4-in.) diameter internal balance. The upper aft balance pin hole 

was used for this installation. 

Pressures over the model wing, aileron, and deflected high-lift system were 

measured by 12 48-S-type scanivalves arranged in two 6-pat modules mounted 

in the fuselage nose. Access to the scanivalves was obtained by removing 

the nose and forward constant sections of the fuselage. In addition to the 

four complete rows of pressure orifices, one pressure tap was located at the 

trailing edge of an inboard station (18-percent semispan) to help evaluate 

any separation that may have occurred (Figure 33). 

The angle of attack of the fuselage reference plane was measured using a 

bubble pack installed in the fuselage nose. From aFRP = -6O to Do, 

the model was pitched using the external pitch system. From O" to +lO" 

angle of attack, the fuselage was pitched using the fuselage internal pitch 

drive while maintaining the balance at O". For angles of attack +lO" to 

+34O, the fuselage was pitched using the external pitch drive with a loo 

angle maintained between the balance axis and the fuselage axis. 

The horizontal stabilizer incorporated remote drive and a 

position-indication system. A Douglas control panel and digital readout was 

provided for use in the tunnel control room. 

LB-507A MODEL INSTALLATION 

The model was mounted in the Ames 12-Foot Pressure Tunnel using the Douglas 

tandem support system and the Ames Task Mark II 10.16-cm (4-in.) balance. 

The balance was attached to the support struts using the Douglas balance 

pitch block. The installation is depicted in Figure 34. 

65 
MODEL LB507A 

I % b/2 1 WING PANEL NOTES I 

72.5% M 

18% 

I 

I 

I (L) 

I 

I (L) 

I 

T.E. ONLY 

FIGURE 33. PRESSURE ROW LOCATIONS 

i 
MODEL LB-507A 

OUTER HATCH 1 

c INNER ACCESS HATCH 

\ I I / 

I BAYONET-. 

PITCH CONTROL 

HATCH -H 

-TURN+ABLE-/--- 

/ STRUT LENGTH 

-iDJUSTMENT 

FIGURE 34. INSTALLATION IN NASA AMES 12-FOOT PRESSURE TUNNEL 

LB-507A RESULTS AND DISCIISSIONS 

Cruise Wing Characteristics 

The initial configuration tested was the cruise wing body with the nacelles, 

pylons, and strakes attached. The basic high-Reynolds-number 

characteristics (lift, pitching moment, and drag) for the configuration are 

shown in Figure 35. Two different runs of the same configuration are shown 

to indicate the repeatability'of the data. This figure indicates that a 

tail-off CLMAX of 1.59 was obtained at the basic test condition of 

M = 0.20 and RNMAC = 4.61 million. This comparedwith a maximum value of 

1.54 obtained from Phase I testing of the LB-486 model. A direct comparison 

of the data from the two tests is shown in Figure 36. Besides a higher 

'LMAX' the LB-507A model exhibited better tail-on pitching moments than 

did the LB-486 model. Though improved, the pitching moments of the LB-507A 

model still included pitch-up prior to stall. Post-stall pitch-down was 

abrupt and forceful. 

67 
- 

MODEL LB-507A ICONFIGURATION B 313W1BX1ElP1CN1CS11F 2 

MACH = 0.20 =1 0. 300 

RN MAC = 4.61 x IO= tl 

e 

8 

5 0.200 

2 

e 

I: 

5 0. 100 

6 

h 

I I .dem 

-10 -5 . 

a 

El -0. It00 

0 

0 

0 

m 

-0.500 

-0. boo 

-0. 700 

q 

J 

RNGLE OF RTTRCK-DEG 

-0. 50 

-0. 800 

STC 

A. LIFT AND PITCHING MOMENT 

Y 

RNGLE OF RTTRCK-DEE 

FIGURE 35. CRUISE WING CHARACTERISTICS AND REPEATABILITY 
Lt. 0 

3. 5 

: 

. 

; 

J 

: 

, 3.0 

L 

; 

: 

: 2.5 

I 

2. 0 

I. 5 

I. 0 

0. 5 I0. DRAG 

FIGURE 35. CRUISE WING CHARACTERISTICS AND REPEATABILITY 
MODEL LB-507 A 

AND 

MODEL LB-486 A 

MACH = 0.20 

RN MAC = 4.61 x IO6 

NACELLES, PY LONS, STRAKES ON 

GEAR OFF t 0.200 -I 0 0 

v 

a I RNGLE OF RTTRCK-DEG - 

I 1 I I I 

$ 

5 IO I5 20 25 

B RNGLE OF ATTRCK-DEG 

-0.502 

WC 

0 -0.200 

-0. 300 

1 

-0.800 

V 

El 

q 

A. LIFT AND PITCHING MOMENT 

FIGURE 36. COMPARISON OF LB-507A AND LB-486B CRUISE WING CHARACTERISTICS 
Y. o3. 5- 

x 

4 

L 

2 

$ ‘3. o- L 

2 

3 

5 ‘2. 5- 

w 

2. oI. 5- 

I. o0.5 

MODEkLLE507A 

MODEL LB-466A 

El 

0 

00 V 0 

0 QQO 

Q B 

0 

0 

El 

v 

SYM RUN TEST ‘H 

0 3 LB-507A OFF 

0 

I I I I I 1 I I I I I I I I I 

0. 02 0. 0) 0. ob 0. 08 0. IO 0. I2 0. 19 0. lb 0.18 o.i.3 0. 22 0. a 0. 2b 0. 28 0. K, 0. 32 0.39 a 

STC DRIZ COEFFICIENT 

0. DRAG 

FIGURE 36. COMPARISON OF LB-507A AND LB486A CRUISE WING CHARACTERISTICS 
Mini-tuft pictures of the wing, for a Mach number of 0.20, are presented in 

Figure 37 for angles of attack before and after C LMAX' 

This figure 

illustrates the stall phenomena of this high-aspect-ratio wing at 

R%AC = 4.61 milli.on. As was the case with the LB-486 model, the outboard 

wing panel stalled prior to the inboard panel. The inboard panel stalled 

completely (separated to the leading edge) at an angle approximately 6O 

higher than the outboard stall angle. Figure 38 presents the chordwise 

pressure distributions of the four streamwise pressure rows for aCLMAX 

(13.61'), and lo and 3' past CYC LMAX* At "'LMAX' suction peaks 

are evident for all spanwise locations. Slightly negative trailing-edge 

pressure coefficients are noted for this condition at all spanwise 

stations. Large spanwise flow angles are indicated in the corresponding 

tuft photo for the trailing-edge region. AtaFRp = 14.59O (lo past 

stall), the 72.5-percent semispan station plot indicates separation near the 

leading edge. At aFRp = 16.54' (2O past stall), the 57-, 72.5-, and 

95-percent semispan stations are separated at the leading edge. On the 

other hand, the inboard station was still heavily loaded. 

Reynolds number and Mach number effects.- The cruise wing configuration was 

also tested at Mach = 0.20 at various reference chord Reynolds numbers, 

ranging from 1.14 million (atmospheric conditions for the Ames facility) to 

4.61 million. Test results are presented in Figure 39. Comparing the 

results of the lowest Reynolds number run to the highest Reynolds number 

data shows that C 

LMAX 

was reduced from 1.59 to 1.31, o~C 

reduced from 14.5' to 13.6', 

LMAX was 

and the magnitude of the post-stall lift 

loss is decreased. A positive CM shift was apparent for angles of attack 

prior to stall, but the configuration still exhibited the same pitch 

variations for the angles just after C 

LMAx' 

The maximum value of L/D was 

reduced from 20.02 to 15.62 by the decrease in Reynolds number. Figure 39 

suggests that C 

LMAX 

will not increase significantly, due to Reynolds 

number effects, as the Reynolds number is increased from the highest wind 

tunnel value to flight conditions. 

72 
A. aFRP = 12.66” 

” aFRP = 13.61° (a 

% ’ MAX 

FIGURE 37. MINI-TUFT PHOTOS-FOR CRUISE WING/EiODY WITH NACE’LLES (RUN 3) (CONTINUED) 

73 
c. Q FRP = 14.59O 

D. aFRp = 16.64o 

FIGURE 37. MINI-TUFT ~0~0s FOR CRUISE wiNG/~00Y WITH NACELLES (CONTINUED) 

74 
75 
MODEL LB-507A 

-II- 

-II- 

-II- 

-10. 

-9. 

-.- 

-7- 

0” -6. 

-5- 

-4- 

-3- 

-2- 

-IPERCENT SEMISPAN = 35.09 

I : I 

6 0 10 12 

d&D 

L. II 20 

PERCENT SEMISPAN = 72.50 

-II 

PERCENT SEMISPAN = 57.00 

-II- 

-12- 

-,I- 

-IO- 

-P- 

-I- 

-,- 

0" -6- 

-5. 

-4. 

-3- 

-2- 

-I - 

0 

I 

!P 

-I3- 

-12. 

-II - 

-IO- 

-9- 

-8- 

-7. 

u" -6. 

-5. 

-4- 

-3- 

-2- 

-1 - 

0 

.iYti RUN MACH ALPHA 

0 =3 0.20 13.61 

, 

I 18 

CHO; 

1) 22 

PERCENT SEMISPAN = 95.00 

I ) I 

22 7 

24 18 26 

CHORD 

A. aFRP = 13.61° (aqMAx) 

FIGURE 38. CHORDWISE PRESSURE DISTRIBUTIONS OF CRUISE WING WITH NACELLES, PYLONS, AND 

STRAKES ATTACHED 

76 
-IS- 

-12- 

-II - 

-ID- 

-6- 

-I- 

-,- 

0” -5- 

-5- 

-a- 

-1- 

-2- 

-IPERCENT SEMISPAN = 35.09 

MODEL LBb07A 

PERCENT SEMISPAN = 72.50 PERCENT SEMISPAN = 95.00 

% 

-1X- 

-12- 

-Il- 

-IO- 

-6- 

-6- 

-7- 

0” -5- 

-5- 

-.- 

-1- 

-2- 

-I - 

6 

PERCENT SEMISPAN = 57.00 

SYM RUN MACH ALPHA 

0 =3 0.20 14.59 

-137 

-12- 

-II- 

-ID- 

-Q- 

-a- 

-7- 

0” -6- 

-5- 

-4- 

-I- 

-2- 

-I - 

0 

I ( 

P 

B. aFRP = ‘14.59O 

FIGURE 38. CHORDWISE PRESSURE DISTRIBUTIONS OF CRUISE WING WITH NACELLES, PYLONS, AND 

STRAKES ATTACHED (CONTINUED) 

77 
MODEL LB507A 

-n- 

-II- 

-II- 

-,D- 

-s- 

-B- 

-I- 

& -5- 

-5- 

-t- 

-Y.- 

-2- 

-LPERCENT SEMISPAN = 36.00 

PERCENT SEMISPAN = 72.50 PERCENT SEMISPAN = 95.00 

-1x- 

-12- 

-II - 

-IO- 

-S- 

-B- 

-7- 

4 -5- 

-5- 

-4- 

-3- 

-2- 

-I- 

-13 

-12 

-II 

-10 

-0 

-I 

-7 

0” -5 

-5 

-4 

-1 

-2 

-I 

0 

I 

PERCENT SEMISPAN = 57.00 

SYM RUN MACH ALPHA 

0 =3 0.20 16.54 

CHUG h 2i 

-II- 

-12- 

-II- 

-10- 

-D- 

-a- 

-1- 

& -s- 

-5- 

-,- 

-3- 

-2- 

-1 - 

C. aFRP = 16.54~ 

FIGURE 38. CHORDWISE PRESSURE DISTRIBUTIONS OF CRUISE WING WITH NACELLES, PYLONS, AND 

STRAKES ATTACHED (CONCLUDED) 

78 
MODEL LB-507A 

CONFIGURATION BgB W, B X,, P,= N,= 

MACH = 0.20 

1.6 

0.8 

1 2 5 10 20 50 

RN MAC ’ lo6 

FIGURE 39. EFFECT OF REYNOLDS NUMBER ON MAXIMUM LIFT OF CRUISE WING 

79 
Figure 40 presents the influence of Mach number on the same configuration. 

These data were obtained at a reference chord Reynolds number of 2.60 

million. The effect of Mach number was to decrease C LMAX ICL~AX= 1044y 

1.40, and 1.34 at Mach = 0.20, 0.26, and 0.32, respectively). Also, 

increased Mach number tended to decrease the angle of attack for the 

outboard stall. 

Macelle/pylon/strake effects.- The effects of the nacelles, pylons, and 

strakes are shown in Figure 41. Removal of the nacelles and pylons resulted 

in a decrease in C 

LMAX 

from 1.59 to 1.47. The pitching-moment curves show 

the nacelles and pylons to be destabilizing prior to stall and stabilizing 

after stall. The drag increment at l.2Vs due to the nacelles, pylons, and 

strakes was 0.0171 and they reduced the L/D from 20.3 to 16.3. Mini-tuft 

photos for the nacelles-off and pylons-off case are shown in Figure 42. 

Be1 ow CLMAXy improvements in local flow, compared to the configuration 

with nacelles, were evident aft of the nacelle location. Outboard 

separation patterns were similar for the nacelles on and off cases; however, 

comparison of Figures 42 and 37 show that the presence of the nacelles 

retarded flow separation on the wing region aft of the nacelles. 

Chordwise pressure distributions for the configuration with the nacelles and 

pylons removed are presented in Figure 43. The angles of attack selected 

are stall (11.55') and higher. At the aFR,, of 13.55', the 

72.5-percent semispan station shows a collapse of the suction peak, while 

the 95-percent semispan station shows only a modest increase in Cpmin and 

mild trailing-edge separation. At a lo higher angle of attack, the 

suction peak of the 57-percent semispan station collapsed. The most 

outboard station remains reasonably well attached up to 16.5O angle of 

attack, the same angle as the nacelles on case. 

From the standpoint of low-speed clean-wing characteristics, the addition of 

strakes to the nacelles is detrimental from both a lift and pitching-moment 

standpoint. This detriment is illustrated in Figure 44. Addition of the 

strakes reduced the tail-off clean-wing C LMAX from 1.62 to 1.59 and 

increased the pre-stall nose-up moments. 

80 
W 

-4 

MODEL LB-507A 

CONFIGURATIONB W X P N S 38 1B 1B 1C 1C 11F 

RN MPC = 2.60 x 106 

we l&l Q 0 

I- 0 0 E 0 

E H 0.300- 0 

2. 50 1

s 

2. oow 

:: 

k 

8 I. 50- 

t 

!I 

I. ooI 

-5a * ) 

I I I I I 

5 IO I5 20 25 

ii 

-0. 100-i 

a 

B -0. 200 

-0. 300 

-0. YOO 

-0.500 

-0. boo 

RNGLE OF RTTRCK-DEG 

0 J RNGLE OF RTTRCK-DEG 

-0.50 

-0. 800 

STC 

A. LIFT AND PITCHING MOMENT 

FIGURE 40. EFFECT OF MACH NUMBER ON CRUISE WING 
w 

ru 

I. o0. 5 

MODEL LB-507A 

0 

0 

q 

0 

+oI mh 

. 4. I I I 1 I I I I I I I I 

12 0. 00 

t I I 

0. 02 0. 

I 

O$ 0. Ob 0. OR 0. IO 0. I2 0. IY 0. lb 0. I8 0. 20 0. 22 0. 2$ 0. 2b 0. 28 0. 30 0. 32 0. 3$ 0. 3b 

STC DRAG COEFFICIENT 

B. DRAG 

FIGURE 40. EFFECT OF MACH NUMBER ON CRUISE WING 
3. 50 

3.00 

2. 50 

% 2. 00 

:: 

# 

k 

$ 

u I. 50 

II. 00 

MODEL LB-507A 

CONFIGURATION B3B W,, XIB 5 

ki MACH = 0.20 

MAC = 4.61 x lo6 

z 0. 30( 

RN t 

8 

u 

5 0. 20c lY 

F 

4 :: 0. IOC l- 0 

l-l 0 

I3 

2 

El 

0 El (L Q -a 

I 1 “-88e 

-10 -5 * 

l-- 

0 u 

3 

-o.ooc ,- 

0 

I3 0 

0 -0. 2oc Iu I 

fi g u5 

I I I I I 

IO I5 20 25 30 

RNGLE OF RTTRCK-DEE 

-0. boo 

I I 1 r I 

5 IO I5 20 25 

-0.700 

RNGLE OF RTTRCK-DEG 

-0. 800 

-0. 300 I- 

-0.YOO I- 

-0. 500 

IA. LIFT AND PITCHING MOMENT 

FIGURE 41. EFFECT OF NACELLES AND PYLON ON CRUISE WING 
MODEL LB-507A 

0 

2. 0 

I. 5 

1 

El 0 I3 I3 

0 

El 0 

0 

0 

- 0. 00 

T I I I I 1 1 I I I I I I I I I I 

,02 0. 02 0. OY 0. Ob 0. 08 0. IO 0. 12 0. IY 0. lb 0. IX 0. 20 0. 22 0. 2Y 0. 2b 0. 28 0. 30 0. 32 0. 39 

STC DRRG COEFFICIENT 

B. DRAG 

FIGURE 41. EFFECT OF NACELLES AND PYLON ON CRUISE WING 
A. aFRp = 12.54’ 

FIGURE 42. MINI-TUFT PHOTOS FOR CRUISE WING/B~DY (RUN 113) 

85 
‘%iGbRE-42. MINI-TUFT PHOTOS FOR CRUISE WING/BODY (CONCLUDED) 
MODEL LB-507A 

-1l- 

-II- 

-II- 

-IO- 

-s- 

-I- 

?- 

E -c- 

-5- 

-a- 

-I- 

-2- 

-1 - 

PERCENT SEMISPAN = 35.00 

PERCENT SEMISPAN = 72.60 

-I3 

-12 

-II- 

-12- 

-II- 

-10- 

-5- 

-.- 

-7- 

0” -5- 

-5- 

-4- 

-,- 

PERCENT SEMISPAN = 57.00 

SYM RUN MACH 

CHDRD 

PERCENT SEMISPAN = 95.00 

A. apRP = 11.55’ (afgMAx) 

FIGURE 43. CHORDWISE PRESSURE DISTRIBUTIONS OF CRUISE WING 

WITHOUT NACELLES AND PYLONS 

87 
PERCENT SEMISPAN = 35.00 

MODEL LB-507A 

PERCENT SEMISPAN = 57.00 

-13- 

-12- 

-II - 

-IO- 

-P- 

-a- 

-7- 

u” -6- 

-5- 

-k- 

-3- 

-2- 

-I - 

PERCENT SEMISPAN = 72.50 

-I,- 

-12- 

-I,- 

-IO- 

-o- 

-I- 

-7- 

-13- 

-12- 

-II- 

-lO- 

-O- 

-5- 

-7- 

2 -6- 

-5- 

-,- 

-3- 

SYM RUN MACH ALPHA 

0 =113 0.20 13.55 

PERCENT SEMISPAN = 95.00 

-13-l 

-1- 

-2- 

B. aFRP = 13.55O 

FIGURE 43. CHORDWISE PRESSURE DISTRIBUTIONS OF CRUISE WING 

WITHOUT NACELLES AND PYLONS (CONTINUED) 

88 
MODEL LB-507A 

PERCENT SEMISPAN = 35.00 PERCENT SEMISPAN = 57.00 

t SYM RUN MACH ALPHA 1 

-Is7 

-II- 

-,,- 

-IO- 

-n- 

-.- 

-7- 

& -5- 

-6- 

-1- 

-I- 

-2- 

-I - 

-13- 

-II- 

-II- 

-IO- 

-6- 

-c- 

-7- 

0” -b- 

-5, 

PERCENT SEMISPAN = 72.60 PERCENT SEMISPAN = 95.00 

-,- 

-3- 

-2- 

D 

1 p,/*- 

,-+------ -.a-.- 

I , 

I6 I. 2d P 24 

CHORD 

-13- 

-12- 

-II- 

-m- 

-6- 

-6- 

-7- 

0” -6- 

-5- 

-‘- 

-5- 

-2- 

-l0 

1 

I2 

-13- 

-12- 

-II - 

-IO- 

-6- 

-c- 

-7- 

0” -6- 

0 =113 0.20 14.50 

) 

P 

____ o- ____ o-------v---u 

1 I4 I6 

Cd 

20 22 

C. aFRP = 14.50~ 

FIGURE 43. CHORDWISE PRESSURE DISTRIBUTIONS OF CRUISE WING 

WITHOUT NACELLES AND PYLONS ATTACHED (CONTINUED) 

89 
MODEL LB-507A 

PERCENT SEMISPAN = 36.00 

-11- 

-12- 

-II - 

-lD- 

-B- 

-I- 

-7- 

OQ -I- 

-5- 

-*- 

-1- 

-2- 

-L- 

-II 

-12 

-II 

-IO 

-D 

-I 

-7 

& -6 

-5 

-4 

-3 

‘1 

PERCENT SEMISPAN = 57.00 

-IS- 

-12- 

-II- 

-*0- 

-9- 

-I- 

-7- 

0” -6- 

-5- 

-a0 1 

---. .__- 

a-------- - 

-__w __*’ 

I , I.0 I 

I2 I. IS 

CHA 

P P 

PERCENT SEMISPAN = 72.50 PERCENT SEMISPAN = 95.00 

D. aFRP = 16.50° 

FIGURE 43. CHORDWISE PRESSURE DISTRIBUTIONS OF CRUISE WING 

WITHOUT NACELLES AND PYLONS (CONCLUDED) 

90 
0 0 

1 y88, w I 

CONFIGURATION B,, W,, X,B PIc N,C S,,, I- n q 

MACH = 0.20 5 

@ El 0 

El 

RN = 4.61 x lo6 tl l-4 0.300- 0 

MAC 

k 0 jTJ0 

El 

ki 

08 El I3 

3.50- 

2 0. 200-, 0 El 

0: 

El 

E 

I: 

I @El 

3. oo- R 0. IOOJ .m 

E 

I? 

ati@ 

I =zeee 

B 

I ’ , 

2. !io- -10 -5 :yl I ’ I I I 1 

1 5 IO I5 20 25 30 

-0. 100 

0 

m 

a -0. 200 

/ 

-I 

; 

2. 00 

# 

:: 

e 

:: 

u I. 50 

RNGLE OF RTTRCK-DEG 

i 

q 

I 

-50 . 

Ge111 

-0. 50- 

STC 

-0. YOO 

-0. 500, 

-0. bO0. 

I I I I I 

5 IO I5 20 25 

-0. 700. 

RNGLE OF RTTRCK-DEG 

-0. 800. 

A. LIFT AND PITCHING MOMENT 

FIGURE 44. EFFECT OF STRAKES ON CRUISE 
Lt. 0 

3. 5 

:: 

h 

5 

K 

=I 3.0 

k 

! 

L 

2 2.5 

2. 0 

I. 5 

I. 0 

0. 5 

MODEL LB-507A 

El 0 0 

El 

El El 

I I I I I I I I 1 I I I I I I I 0. 09 0. Oh 0. ox 0. IO 0. 12 0. Ilt 0. lb 0. IX 0. 20 0. 22 0. 2Y 0. 2b 0. 28 0. 30 0. 32 0. 3Y 0.' 

STC DRRG COEFFICIENT 

B. DRAG 

FIGURE 44. EFFECT OF STRAKES ON CRUISE 
Landing Configuration Characteristics 

The primary landing configuration consisted of: 

1. a two-segment flap deflected at Z"/lZo (main flap/auxiliary 

flap) 

2. a slotted, leading edge, outboard slat deflected at 27' 

3. a slotted slat or short-chord FCK inboard. 

The grid optimization studies for the inboard slat and inboard FCK are shown 

in Figures 45 and 46, respectively. A comparison of the best slat position 

versus the two best FCK positions is presented in Figure 47. The best 

pitching-moment characteristics were those associated with the FCK deflected 

at 70°. This configuration also resulted in the highest tail-off C 

Deflecting the FCK at 55' decreased the 

LMAX 

of the test, 3.08. C 

LMAX 

from 

3.08 to 2.94, decreased the stall angle from 17.2' to 15.Z", and 

degraded the post-stall pitching moments. The inboard slat configuration 

had a 'LMAX value between the two FCK values and exhibited the most 

undesirable pitching moment trends of the group. 

Reynolds number and Mach effects.- The effect of Reynolds number on the 

maximum-lift coefficient of the landing FCK configuration is shown in 

Figure 48. Unlike the trends for the cruise wing, the trends for the 

landing configuration suggest that the C 

LMAX 

of the landing configuration 

will increase beyond the wind tunnel values as the Reynolds number is 

increased from the highest wind tunnel value to flight Reynolds number. Any 

effort to extrapolate the data to arrive at an estimated C 

LMAX 

value for 

flight conditions would be unwise in light of the distinct break in the 

cL MAX 

versus Reynolds number curve for the cruise wing (Figure 39). 

The effect of Mach number on the maximum-lift coefficient for the same 

landing configuration is depicted in Figure 49. Again the trends of the 

cruise wing differed slightly from those of the landing configuration. 

Whereas C 

LMAX 

of the cruise wing decreased monotonically with Mach number, 

the 'LMAX of the landing configuration increased slightly as the Mach 

number was increased from 0.20 to 0.26. As the Mach number was further 

increased to 0.32, the C 

LMAX 

of the landing configuration decreased from 

2.88 to 2.79. 

93 
P 1 

3. 50. 

3. co 

2. 50. 

I- 2. 00. 

z 

=1 

r L 

i I. 50. 

t 3 !!! . 

Q 1.00, 

Q 

8 

0. 50. 

I 

-5 * 

$8 

-0. 50. 

STC 

CONFIGURATION B ~AWl~XIBPICNICSllF 

MACH = 0.20 

RN MAC = 4.61 x IO6 

6 FLAP = 25112.5 

-0. 400 

-0. boo 

e 

I 1 1 1 I 

5 IO 15 20 25 El 

q go. A0 

RNGLE OF RTTRCK-DEG 

,;A- A 

v q 

I I I I I 

5 IO 

I52 2o 25 3o 

V El 

0 El A 

RNGLE OF RTTRCK-D&J 

0 

INBOARD 

6 SLAT 

128 

12A 

9A 

8B 

128 

12c 

RUN 

17 I 19 

20 

21 

22 

23 

A. LIFT AND PITCHING MOMENT 

FIGURE 45. LANDING SLAT GRID OPTIMIZATION 
9. 0 

8. 0 

7. 0 

$ 

. 

i 

t: 

j b.0 

: 

;’ J 

: 5.0 

* 

Y. 0 

3. 0 

2. 0 

I. 0 

- 

St 0 

STC 

‘. 0. 

MODEL LB-507A 

vV V 

B 

A 

A 

A 

0 

0 

V 

4 V 

qfr OA v 

-4 

OUTBOARD 

6 SLAT 

278 

27A 

27A 

27A 

27A 

27A 

I I I I 1 I 8 I I I I I I I I I i 

0 0. 04 0. 08 0. 12 0. lb 0. 20 0. 2Y 0. 28 0. 32 0. 3b 0. 90 0. LtY 0. 48 0. 52 0. 5b 0. b0 0. b’t 0. bX 

DRAG COEFFICIENT 

B. DRAG 

FIGURE 45. LANDING SLAT GRID OPTIMIZATION 
8 

3. 50 

3. 00 

2. 50 

0. 50 

I 

-5 . 

66 

-0. 50 

STC 

-. 

MODEL LB-507A 

CONFIGURATION B 3BW,BX1BN1CPIC811P 

MACH = 0.20 

RN MAC = 4.61 x lo6 

6 = 25/12.5 

FLAP 

6 = XXl27A 

LE 

i 

# 

F 0. 300, 

k 

B 

5 0. 200, 

!k 

g 

1 I I , I 

5 IO 15 20 25 

RNGLE OF RTTRCK-DEG 

-0. bO0, 

f! 60. 700. * 

-0. 800. 

A. LIFT AND PITCHING MOMENT 

FIGURE 46. FCK OPTIMIZATION WITH LANDING FLAPS 
9. 0 

8. 0 

MODEL LB-507A 

-- 

I I I I I I I I , 8 I I I I I I IO 0. o+ 0. ox 0. I2 0. lb 0. 20 0. 2Lt 0. 2x 0. 32 0. 3b 0. YO 0. LtY 0. 98 0. 52 0. 5b 0. b0 0. b’t 0. b8 

DRRG COEFFICIENT 

B. DRAG 

FIGURE 46. FCK OPTIMIZATION WITH LANDING FLAPS 

72 
3. 50 

3. 00 

2. 50 

-0. 50 

STC 

MODEL LB-607A 

CONFIGURATION B 3BW113XfBP1CNICS11F 

MACH = 0.20 

RN MAC = 4.61 x lo6 

6 FLAP = 25112.6 

0 

I? 

w 

I 0 I 

m ANGLE OF 

-0. 300- 

-0. Ltoo- 2 

0 

@f@ 

-0.500- w” 

wo 

90 

10 

I I 

5 IO 

RNGLE OF ~~TTRCK-DEG 

1 

A. LIFT AND PITCHING MOMENT 

FIGURE 47. FCK AND SLAT COMPARISON WITH LANDING FLAPS 
8 

9. 0 

8. 0 

7. 0 

I 

i 

i 

kj b.0 

2 

4 

5 5.0, 

v 

Y. 0, 

3. 0, 

2. 0 

I. o8 

Q!l 

MODEL LB-607A 

0 

El 

0 

q 

0 

q 0 

El 0 

--ec, 

* 0.00 

I I I I I I I I I 1 I 1 I I I I I 

OY 0. OY 0. OS 0. 12 0. lb 0. 20 0. 2Y 0. 2X 0. 32 0. 3b 0. YO 0. LtY 0. Y8 0. 52 0. 54 0, b0 0. b’t 0. b8 0 

STC DRRG COEFFICIENT 

B. DRAG 

FIGURE 47. FCK AND SLAT COMPARISON WITH LANDING FLAPS 

72 
MODEL LB-507A 

CONFIGURATION B 313W1BX1BP1CN1C 

MACH = 0.20 

6 LE 

= 70Al27A 

6 FLAP = 25112.5 

2.4 

3.2 

2.8 

2.6 

0 

0 

0 

2 5 10 20 50 

RN x10 6 

MAC 

FIGURE 48. EFFECT OF REYNOLDS NUMBER ON MAXIMUM LIFT OF LANDING FCK/SLAT 

CONFIGURATION 

100 
MODEL LB-507A 

CONFIGURATION B W 3B 1A ‘IS ‘IC NlC GIA 

RN MAC = 2.89x106 

6 LE = 70127A 

6 FLAP = 25il2.5 

TAIL OFF, GEAR DOWN 

2.4 I I I I I 

0.20 0.22 0.24 0.26 0.28 0.30 0.32 

MACH NUMBER 

FIGURE 49. EFFECT OF MACH NUMBER ON C 

LMAX 

LANDING FCK/SLAT CONFIGURATION 

Nacelles/pylons/strakes effect.- Figure 5J shows the effects of having the 

nacelles, pylons, and strakes on the landing configuration with the slat 

inboard. The nacelles and pylons had a degrading effect on the post-stall 

pitching moments in that their addition eliminated the post-stall pitch-down 

that was present (tail-off) with the nacelles and pylons off. The nacelles 

and pylons had no significant effect on the maximum lift value for this 

particular configuration. 

The nacelle strakes, which were added to increase the C 

LMAX 

of the inboard 

slat configurations, were effective in that respect. The tail-on data of 

Figure 51 showed that the strakes increased the tail-on CLMAX of the 

inboard slat configuration from 2.94 to 3.08. As might be expected, the 

strakes degraded the pitching-moment characteristics. Figure 52 shows that 

the strakes had very little impact on the inboard FCK configuration. 

101 
3. 50 

3. 00 

2. 50 

!i 2. 00 

E 

E 

2 

S I. 50 

5 8 

Jg 

I. 00 

El 

0 

0. 50 

I . ee 

-5 

-0. 50 

STC 

0 

J 

MODEL LB-507A 

CONFIGURATION BsB W,, Xle 

IB 

MACH = 0.20 =1 RN = 4.61 x lo6 z 0.300- 

M’AC 

6 LE = lZCl27A 

B 

s 

6 = FLAP 25i12.5 g 0. 200- 0 

it 

8 0 0 

l!! 0 

0 E ?.I 0. IOO- 00 

El 0 Q ti 0 

Q a 8 0 I 

&IO 

1 ve@J I I I 1 I 

q -5 * Q 5 IO I5 20 25 30 

0 

0 

-0. IOO- 

-0. zoo0 

El 

I I I I I 

5 IO 15 20 25 

RNGLE OF ATTRCK-DEG 

El 

RNGLE OF RTTRCK-DEE El 

0 q El El 

q 

I 0 El 

-0. YOO- 

-0. 500- 

A. LIFT AND PITCHING MOMENT 

-0. 300 

1 

0 B 

0 El 

-0. bOO- 

@ 

-0. Eo- I 

R 

-0.800’ 

B 

El 

FIGURE 50. EFFECT OF NACELLES AND PYLONS ON LANDING SLAT CONFIGURATION 
C. 

9. c 

8. C 

7. 0 

E 

h 

; 

i 

$ b.0 

L 

5 

2 

3 5.0 

.d 

Y. 0 

3. 0 

2. 0 

I. 0 

'1 

‘1 

-I 

MODEL LB-507A 

0 

El El 

0 

Yt .0. 00 0. OLt 0. 08 0. 12 0. lb 0. 20 0. 2+ 0. 28 0. 32 0. 3b 0. YO 0. Yt 0. Y8 0. 52 0. 5b 0. b0 0. b’t 0. b8 ( 

STC DRRG COEFFICIENT 

B. DRAG 

FIGURE 50. EFFECT OF NACELLES AND PYLONS ON LANDING SLAT CONFIGURATION 
3. 50 

3. 00 

2 50 

Q 1.00 

B 

0. 50 

MODEL LB-507A 

CONFIGURATION B W 38 ,BX,BPtCN,CS,lFVIDHID 

s W 

MACH, = 0.20 =I 

RN MAC = 4.6’1 x lo6 

t” 0.300- 

b 

6 LE = lZCl27A e 

61 

m 

0 

Q 

m 

I 

g 

0 

1 

-5 9 5 IO 15 20 25 

J RNGLE OF RTTRCK-DEG 

-0. 50 

STC 

RNGLE OF RTTRCK-DEG 

OQ 

-0. 200 

A I 

-0. bOO;l 

-0.7ooJ 

I 

-0. 800; 

A. LIFT AND PITCHING MOMENT 

J 

FIGURE 51. EFFECT OF STRAKES ON LANDING SLAT CONFIGURATION 
9. 0 

8. 0 

7. 0 

b 

. 

i 

tl 

: b.0 

i 

i 

: 

3 5.0 

I 

Y. 0 

3.0 

2. 0, 

I. 0, 

- 

BY 0 

STC 

MODEL LB-507A 

El 

0 

II 

0 

El 

8 

B 

B 

B 

B 

I I I I I I I I I I I I I I I I I 

IO 0. OY 0. 08 0. I2 0. lb 0. 20 0. 2+ 0. 28 0. 32 0. 3b 0. 40 0. w 0. 48 0. 52 0. 5b 0. b0 0. b’t 0. b8 I 

DRRG COEFFICIENT 

B. DRAG 

FIGURE 51. EFFECT OF STRAKES ON LANDING SLAT CONFIGURATION 

1 

72 
-- 

MODEL LB-507A 

+w 

CONFIGURATION B 38 lBXIBPICNICS1lFVIDHID 

W Li 

k! 

3.50 

3. 00, 

L q 

-I 

0 1.00 

I 

0. 50 

I 

-5 - 

88 

-0. 50. 

STC 

MACH = 0.20 

RN MAC = 4.61 x 106 

6 

LE = 70Al27A 

6 

FLAP = 25/12.5 

lil 

; 0. 300 

e 

:: 

ii 0.200 F 

z 

5 #q 0. 100 

B 

i 

-10 I -5 I ~~ . c 5 II I , I I 

IO 15 20 25 30 

-0. 100-j B I 

-0. 200 

00. 300 

B 1 

RNGLE OF RTTRCK-DEG 

-0. bO0 

I I I I 

5 

1 

IO I5 20 

25 

-o.700 

RNGLE OF RTTRCK-DEG 

-0. 800 1 

A. LIFT AND PITCHING MOMENT 

FIGURE 52. EFFECT OF STRAKES ON LANDING FCK CONFIGURATION 

1 
v. 0 - 

MODEL LB-507A 0 

B 0 

0 

8. 0 

7. 0 

I 

. 

; 

rl 

: 

, b.0 

i 

: 

: 

: 5.0 

r 

8 

0 

8 

0 

a 

It. 0 

3. 0 

2. 0 

I. 0 

*. TN 0. 00 0. OY 0. 08 0. 12 0. lb 0. 20 0. 2Y 0. 28 0. 32 0. 3b 0. YO 0. YLt 0. YE 0. 52 0. 5h 0. ho 0. b’t o.bX 0 

STC DRAG COEFFICIENT 

B. DRAG 

FIGURE 52. EFFECT OF STRAKES ON LANDING FCK CONFiGURATiON 
Large inboard flap deflection effect.- In addition to testing the baseline 

landing flap deflection of 25°/12.50, a deflection of 350/10° was 

tested at two different grid positions. The original grid position included 

a negative overhang of l%, and resulted in a slight reduction in C 

LMAX 

from that of the baseline deflection (Figure 53). Analysis of the mini-tuft 

photos (Figure 54) and the trailing-edge press,ures (Figure 55) indicated 

that the large deflection caused separation in the trailing-edge region. In 

order to reduce the extent of trailing-edge separation, a new grid position 

including a positive overhang of 1 percent was created by extending the 

spoiler trailing edge. As the mini-tuft photos and trailing-edge pressures 

show, the positive overhang was effective in reducing trailing-edge 

separation problems. CLa = o increased by nearly 0.20 and C 

increased compared to the baseline but only by 0.03. 

LM.AX 

The large deflection 

did, however, result in a large drag increment at 1.3Vs (0.0405 and 0.0270 

for the negative and positive overhang cases, respectively). 

Takeoff Configuration Characteristics 

Most of the work accomplished with takeoff configurations was directed 

toward the use of sealed (zero gap) slats. The advantage of the sealed slat 

is that it results in appreciably higher L/D values. The disadvantages are 

that it provides lower values of CLMAX and can result in poor stalling 

characteristics, particularly if a small amount of yaw is present at stall. 

Figure 56 compares data for the slotted and sealed outboard slats,with an 

FCK deflected at 55O inboard. The slat grid 20A was completely sealed, 

the grid 208 had a small gap, and the grid 27A had a normal gap. As the gap 

was decreased, the tail-off C LMAX 

decreased from 2.55 to 2.40 and the 

pitching moments became more positive. The L/D values at 1.2Vs, on the 

other hand, increased from 11.97 to 12.87. The mini-tuft photographs of 

Figure 57 clearly show the earlier separation of the outboard panel for the 

sealed slat configuration. 

108 
” 

1 

3. 50 

ki 

2. 00 

:: 

5 

b 

I.5 I.90 

0. 50 

r 

-5 * 

38 

-0. 50, I 

STC 

MODEL LB-507A 

CONFIGURATION B W 38 lBXIBPICNICSllF 

t 

E 

MACH = 0.20 

RN MAC = 4.61 x lo6 

i 0.300 

kl 

6 LE = lZCl27A 8 

5 0.200- 0 

L 

: 

2 0. IOO6 99 

i? 0. ii 

*-II0 15 . e6e r 

n 

5 IIO I I5 I - 0 a8 2o I 25 1 1 30 

0 

-0. IOOom 0 

RNELE OF ,TTR,%Ef 

-0. zoo- 

* q * 

El 

00 

-0. 300- go* 

000 

-0. YOO- /y 

0 8 

1 0 w- -0. 500 0 c 

I I I I I 

5 IO I5 20 25 

RNGLE OF RTTRCK-DEG 

El 

I3 -0. 800 

A. LIFT AND PITCHING MOMENT 

FIGURE 53. EFFECT OF LARGE FLAP DEFLECTION 
9. 0, 

8. 0, 

7. 0, 

z 

. 

r 

i 

r: b. 0. L 

; 

3 

: 5. 0, 

,s 

$. 0, 

3. 0 

2. 0 

I. 0 

- 

Q gQ 

0 0 El 

4 I I I I I I I I 1 1 I I I I I I I 

OY 0. 00 0. o+ 0. ox 0. I2 0. lb 0. 20 0. 2Y 0. 28 0. 32 0. 3b 0. +o 0. YY 0. YX 0. 52 0. 5b 0. b0 0. bY 0. bX ( 

STC DRRE COEFFICIENT 

B. DRAG 

FIGURE 53. EFFECT OF LARGE FLAP DEFLECTION 
A. O.H. = 1% i”FRP E 19.12’, 2’ PAL8 

= 20.08”. 4- UCbllAX’ 

FIGURE 54. MINI-TUFT PHOTOS FOR 3s 

0 FLAP DEFLECTION CONFlGURATlONS 

111 
-0.8 

1 

a.6 - 

CF 

TE 

MODEL LBBB7A 

CONFIGURATION B fE W IA ‘,A ‘1C NIC 

M = 020 

RN MAC = 4.61 x lo6 

6 LE = 12Cl27A 

TAIL OFF 

t7 = 0.18 

0 * 0 

0 

I I I I I ~-. mu 

4 8 12 16 20 24 28 

QFRP - DEG 

q o 

00 0 

00 

cP TE 0 O Oo 

-0.4 -0 

1 

0 00 0 0 0 

a 0 

-0.2 -0.2 0 -0 0 qo 000 

0 0 

-0.8 - 

q = 0.35 

-0.6 - 

“FRP “FRP - DEG 

a.2 L - 

FIGURE 55. EFFECT OF LARGERGE FLAP DEFLECTION DEFLECTION ON TRAILING TRAILING EDGE PRESSURES PRESSURES 

1.12 
-4.8 

-0.8 

-0.6 

CF 

TE -0.4 

-0.2 

0 

MDDEL LB-507A 

q = 0.725 

0 

0 

0 

0 

0 0 000 q n 

0 q ooo 3 0 0. 0 

0 q 

E!- 

3 ----kyko 

0 I I 

l2 0 l6 ok 

ffFRP - DEG 

FIGURE 55. EFFECT OF LARGE INBOARD FLAP DEFLECTION ON TRAILING EDGE PRESSURES 

(CONCLUDED) 

113 
3. 50 

3. 00 

2. 50 

-0. 50 

STC 

MODEL LB-507A 0 

CONFIGURATION B W X lAPICNICSllF t 0 

38 16 

MACH = 0.20 

!! 0 El t” 0.300- 0 o 0 

RNinAC = 4.61 x 106 ki 0 I3 

6 = 5/10 Ei 

FLAP 

5 0.200- 

000, 

800 

E go 0 0 

A 

0 

f I3 m 0 ooo(Tj 0 

6 0. IOO- A 

I 

E 

ii! I!3 m 

.01 

, 1. 

b 

I 4 I I 

5 IO & 15 20 A 25 30 

go 

A 

-0: IOO- a 

ATTRCK-D:i? ‘13 ’ 

-0. 200- 

I 1 I I I 

5 IO I5 20 25 

-0. bO0 

-0. 700 

RNGLE OF RTTRCK-DEG 

1 

A. LIFT AND PITCHING MOMENT 

FIGURE 56. OUTBOARD SLAT GRID OPTIMIZATION WITH FCK INBOARD 
9. 0 

8. 0 

Y. 0 

3. 0 

2. 0 

I. 0 

---$;e 

3Y c 

STC 

IrI. b 

MODEL LB-507A 

0 

0 

El 0 

0 

0 0 

0 

8 

m 

27A 

208 

20A 

20A 

lx?%* r? I .- I I 1 I I I I I I I , I I I I 

0 0. OY 0. 08 0. 12 0. lb 0. 20 0. 2Y 0. 28 0. 32 0. 3b 0. YO 0. wt 0. YX 0. 52 0. 5b 0. b0 0. bit 0. bX 0 

DRRG COEFFICIENT 

B. DRAG 

FIGURE 55. OUTBOARD SLAT GRID OPTIMIZATION WITH FCK INBOARD 

72 
A. SLOTTED SLAT OUTBOARD, aFRP = 20.94O 

B. SEALED SLAT OUTBOARD, aFRP = 20.B!Yi” 

FIGURE 57. MINI-TUFT PHOTO OF TAKEOFF CONFIGURATION SHOWING EFFECT OF OUTBOARD 

SEALED SLAT 

-116 
Figure 58 compares the results of a sealed slat outboard with three 

different inboard leading-edge configurations: a slotted FCK, a sealed 

slat, and a clean leading edge. Because of the early stall of the inboard 

wing not protected by a leading-edge device, the C 

LMAX 

of the clean 

configuration was very low (2.09) and the pitching moments were very well 

behaved. The C 

LMAX 

of the inboard sealed-slat configuration was 2.24 

while that of the slotted FCK was 2.40. The pitching-moment trends of the 

FCK and the sealed slat were similar: both showed nose-down moments just 

after stall, even in the absence of a tail. The respective values of L/D at 

1.2V, for the inboard clean leading edge, sealed slat, and FCK are 14.59, 

13.50, and 13.57, respectively. 

One concern with the sealed slats is that they can result in lateral 

instability when stall occurs under a yawed condition. This tendency is 

illustrated in Figure 59. With a sealed slat outboard, the inboard 

sealed-slat configuration became laterally unstable at aFR,, = 19O; the 

FCK at o~,-RR = 17.5'. However, with a slotted slat outboard, the 

FCK/slat configurations remained laterally stable throughout the 

angle-of-attack range investigated (Figure 60). 

Strakes effects.- Figure 61 shows that the addition of nacelle strakes to 

the takeoff configuration with sealed slats inboard and outboard caused only 

small changes in the lift and drag characteristics. The CLMAX increment 

due to the strakes in conjunction with takeoff flaps and slats, 0.06, was 

less than half that for the landing flaps and slats case, 0.14. As was the 

case with clean wing and landing configurations, the strakes were 

detrimental to the pitching-moment characteristics. 

Mach number and Reynolds number effects.- Figures.62 and 63 show the effect 

of Mach number and Reynolds number, respectively, on the aerodynamic 

characteristics of the takeoff configuration with an FCK inboard and a 

sealed slat outboard. As the Mach number was increased from 0.20 to 0.32, 

'LMAX 

decreased from 2.20 to 2.15 and the pitching moments degraded 

slightly. Below CLMAX, the drag polar was insensitive to Mach number. 

The %MAX versus Reynolds number curve of Figure 63 suggests that the 

maximum lift coefficientwilI,continue to increase as the Reynolds number 

increases towards the flight value. 

117 
_. .__ 

MODEL. LB-607A 

CONFIGURATION B 

38 WIB ‘1B ‘1C NIC ‘11, 

5 

:: 

MACH = 0.20 E 0.300 

RN = 4.61 x lo6 k 

MAC w 

cob0 

I 1 -0.500 1 

INBOARD 

SYM RUN LE DEVICE 6 

A. LIFT AND PITCHING MOMENT 

FIGURE 58. EFFECT OF INBOARD LEADING EDGE DEVICE WITH A SEALED SLAT OUTBOARD 
v. o-# MODEL LB-507A 

8. oP 

3 

tt n b. OG 

5 5.0 I -1 ” 

Lt. 0 

3. 0 

2. 0, 

1.0. 

--e+ 

OY 0 

STC 

0 

El 

El 

El 

El I3 El El 

El El 

0 0 

0 

0 

0 

E 

I I I I 1 I I I , I I I I I I I 

10 0. OY 0. 08 0. 12 0. lb 0. 20 0. 2Y 0. 28 0. 32 0. 3b 0. YO 0. 99 0. $8 0. 52 0. 5b 0. b0 0. b’t 0. b8 C 

DnR; COEFFICIENT 

B. DRAG 

FIGURE 58. EFFECT OF INBOARD LEADING EDGE DEVICE WITH A SEALED SLAT OUTBOARD 

72 
;s 

0 

MODEL LB-507A 

CONFIGURATION B 38 W 1B ‘1El ‘IC N,C HIA ‘1, 

MACH = 0.20 

t 

0 

STABLE 

0 

(AC,) 

A/3=-5’ 

RN 

I I I I 

-10 -5 0 5 10 15 30 

(AC”) 

Afl=--5O 

-0.01 - 

I I I I I I I I 

-10 -5 0 5 10 15 20 25 30 

ANGLE OF ATTACK (DEG) 

FIGURE 59. EFFECT OF SLATS ON ROLLING AND YAWING MOMENTS THROUGH STALL 

WITH SIDESLIP 
t 

0.08 

STABLE 

0.06 - 

(AC,) 

Af3= -5O 

0.04 - 

0.02 - 

---- 

rC 

-A-- 

MODEL L0-507A 

CONFlGURATtON B 38 wlB ‘1, ‘IC NIC G,A 

MACH = 0.20 

RN MAC = 4.61 x lo6 

6 FLAP = 25/lie 

I 

-5 

-0.02 l- 

‘La 

LMAX 

I I I I I 

5 10 15 20 25 

ANGLE OF ATTACK (DEG) 

FIGURE 66. EFFECT OF fNBOARD FCK SLAT DEFLECTION ON ROLLING MOMENT THROUGH STALL 

WITH SIDESLIP 

121 
MODEL LB’607A e 

CONFIGURATION B 38 W ,BxlBpICNICsllP”lDHtD 8 

MACH = 0.20 z 0. 300- 

RNMAC = 4.61 x lo6 k 

4: 

3. 50 6 = 6A/20A l-l 

LE 6 = mo z 0. zooFLAP 

‘H . =o E 

3. 00 4 H 0. IOOE 

E- 2. ooI5 

tl 

tl 

kl 

8 I. 50- 

t 

!7 

RNELE OF RTTACK-DEG 

0 

0 

Q 

-0. 500- 

0. 50 q 

0 

-0.0. 700- boo- El 

0 

El 

q 0 0 0 El 

El 

STRAKES SYM RUN El 

-0.5oJ ON 

-0. 800’ OFF 

STC 

A. LIFT AND PITCHING MOMENT 

FIGURE 61. EFFECT OF STRAKES WITH TAKEOFF FLAPS AND SEALED SLATS 
9. 0. 

8. 0. 

7. 0. 

x 

- 

i 

: 

1: b. 0. 

L 

2 

3 

3 5. 0. 

, 

Y. 0. 

3. 0. 

2. 0. 

MODEL LB-507A 

0 

El 

0 

I3 

STC DRRG COEFFICIENT 

B. DRAG 

FIGURE 61. EFFECT OF STRAKES WITH TAKEOFF FLAPS AND SEALED SLATS 
3.50- 

3. oo2.50- 

l.OO- #y@ 

1 

MODEL LB507A c- 8 

CONFIGURATION B 

38 1B 1B lCplCsllF W X N E 

RN 

=I 

= 

MAC 2.60 x IO6 7 

0. 300- m w 

6 = 5/10 e 0 @ 0 0 

FLAP :: 

6 

om 

LE = 55Al20A ‘z 0.200- 

2 

G? l!l 0 

f, 0. IOO6 

h 

I+ 

h r 1 ,eee . I I I 1 1 I 

-10 -5 0 5 IO I5 20 25 30 

9 

e 

0 -0. IOO- El 

VELE OF RTTRCK-DEE 

-0. 200 10 

@ 

b3 -0. 300 Q 

I . 

-0. uo0-l 

1 

SYM RUN MACH ‘,,A, -0. I300 I I I 

015 -;;-0. 700 - 

-0. zooSTC 

A. LIFT AND PITCHING MOMENT 

FIGURE 62. EFFECT OF MACH NUMBER ON TAKEOFF FCK/SLAT CONFIGURATION 
9. 0 MODEL LB-507A 

go 

3. o8 

B 

2. oQ 

I. o- d 

d 

IQ 

4.3 1 I I *, I I , I I I I I I I I I I ---- IY 0. 00 0. OY 0. OX 0. I2 0. lb 0. 20 0. 2Y 0. 28 0. 32 0. 33 0. YO 0. YLt 0. w 0. 52 0. 5b 0. b0 0. bY 0. I38 0 

STC DRRG COEFFICIENT 

6. DRAG 

FIGURE 62. EFFECT OF MACH NUMBER ON TAKEOFF FCK/SLAT CONFIGURATION 

72 
2.8 

2.6 

2.4 

2.0 

i .a 

MODEL LB507A 

CONFIGURATION B 3BW1BX1BP1CN1C 

MACH. = 0.20 

6 

LE = 55Ai20A 

6 FLAP = 6110 

1 2 5 10 20 50 

RN MAC ’ lo6 

FIGURE 63. EFFECT OF REYNOLDS NUMBER ON MAXIMUM LIFT OF TAKEOFF FCK/SLAT 

CONFIGURATION 

126 
Alternative Flap Settings.- In addition to the primary takeoff flap setting 

of 5"/1OO, two other takeoff flap settings (O"/Oo and 150/10°) 

were tested. Figure 64 presents the basic aerodynamic characteristics for 

the 15"/10° flap setting with a variety of leading-edge-device 

combinations. The highest C 

LMAX 

was associated with the slat/slat 

configuration. The best pitching moment was associated with the FCK/slotted 

slat configuration. The highest L/D values were associated with use of a 

sealed slat outboard. 

The basic aerodynamic characteristics of the aircraft with a clean trailing 

edge are presented in Figure 65 for several leading-edge device 

combinations. The combinations investigated included a sealed slat outboard 

with an FCK or sealed slat inboard, and a slotted slat outboard with a clean 

leading edge inboard. This latter configuration was representative of an 

auto-slat system. Also shown are the characteristics of the cruise wing, 

for reference. The pitching-moment curves show the obvious aerodynamic 

benefit of an auto-slat system in improving stall behavior. Figure 66 

summarizes the L/D values for the takeoff configurations. 

Aileron and Spoiler Characteristics 

Aileron effectiveness is presented for takeoff and landing configurations in 

Figures 67 and 68, respectively. At pre-stall angles of attack, the aileron 

effectiveness was well behaved for most angles of attack, but near the stall 

angle the effectiveness of the upward deflected aileron diminished. The 

shape of the rolling moment curve with aileron deflection indicates, for all 

flap settings, that the negative deflections (TEU) were more effective than 

the positive deflections (TED). In many cases, the incremental rolling 

moment obtained was more than twice as large as the corresponding value for 

positive aileron deflection. (Good data for the landing flaps, with 

positive aileron deflections are not available.) 

127 
3. 50 

1 

MODEL LE507A + 

CONFIGURATION B 36 1B lBPICNICSllF W X 6 

MACH = 0.20 z 0. 300 

R%AC = 4.61 x lo6 tl 

6 

FLAP = 15110 

3. 00 -I 5 0. 100 

2. 50 

I 

2. oo- ,Q 

8 

I. 50- 

Q 

8 

I oB 

q 

0 

0 

0. 50- 

I 

-5 * 0 

I I I I I 

5 IO I5 20 25 

RNGLE OF RTTRCK-DEE 

-0. 50- 

STC 

e 

og 

q 

I I I I I 1 

5 IO OkfJ 20 25 30 

08 

w 

CK-DEG 

SYM RUN LE DEVK :E I 6 LE I 

A. LIFT AND PITCHING MOMENT 

FIGURE 54. AERODYNAMIC CHARACTERISTICS OF THE 15°/100 FLAP CONFIGURATIONS 
9. 0 1 MODEL LB-507A 

2. oI. o0 0 

-6CI 0 1JI 

* 0. 00 

I I I I I I I I I # I I , 1 I 

OY 0. OLt 0. ox 0. 12 0. lb 0. 20 0. 2s 0. 28 0. 32 0. 3b 0. YO 0. LtY 0. YX 0. 52 0. 5b 0. b0 0. b4 0. b8 0 

STC DRRE COEFFICIENT 

B. DRAG 

FIGURE 54. AERODYNAMIC CHARACTERISTICS OF THE 15°/100 FLAP CONFIGURATIONS 

72 
MODEL LB-667A 

CONFIGURATION B 38 W 1B lBplC NICSllF X 

MACH = 0.20 

RN MAC = 4.61 x IO6 El 

3.50- TAIL OFF 

q 

Itoo2. 50- 

-0. 800- 

STC 

RNGLE OF RTTRCK-DEG 

A. LIFT AND PITCHING MOMENT 

FIGURE 65. AERODYNAMIC CHARACTERISTICS OF CLEAN TRAILING EDGE CONFIGURATIONS 
- 

. 

.-c _ 

j 2.57 

I 

2. 0’ 

I. 5- 

I. o0. 5- 

MODEL LB-507A 

D 0 

I3 

m 

0 

B q O q B 

q 

C 

0 

L 

@” hh OQ 00 0 

-ec; lTw3 A 

* 

TY 1 I I 1 I I I I I I I I I I I 

D2 0.00 0. 02 0. 09 0. Ob 0. 08 0. IO 0. I2 0. IY 0. lb 0. 18 0. 20 0. 22 0. 29 0. 2b 0. 2X 0. 30 0. 32 0.39 ( 

STC DRRG COEFFICIENT 

B. DRAG 

FIGURE 65. AERODYNAMIC CHARACTERISTICS OF CLEAN TRAILING EDGE CONFIGURATIONS 

1 
18 

8 

6 

4 

MODEL Lp-507A 

CDNFlGlJAAilON 6 

MACH = 0.20 

RNhlAc = 4.61 x lo6 

6 FLAP = o/o 

6 FLAP = 5110 

6 

INBD,%TBD 

LE DEVICE 6 FLAP 

SYMBOL INBDlOUTBD MAIN/AUX 

8°f200 SLAT*ISLAT= o/o. 5ilO. 15flO 

---- 5!i”1200 FCKlSLAT* o/o. 5/10.15/10 

j l * * 12Ol27.5’ SLAT/SLAT om. 5110. xv10 

l -•-•-+ 55Oi27.5’ FCKISLAT 0l0.5/10.15/10 

*SEALED SLAT 

0.4 0.6 0.8 1 .o 1.2 1.4 1.6 1.8 2.0 

cL 

FIGURE 66. TAKEOFF L/D SUMMARY 

132 
MODEL LB-!iO7A 

CONFIGURATION B 

3BW1BX1BPlCNlCVl~H~~ 

MACH = 0.20 

RN MAC = 4.61 x lo6 

6 LE = 8A/20A 

6 FLAP = 5110 

i, = 0 

TED 

8 

4 

0 

16 

Q (DEG) 

b TEU 

-10 -20 

6 aRH (DEG) 

FIGURE 67. ROLLING-MOMENT COEFFICIENT DUE TO AILERON DEFLECTION FOR SEALED SLAT 

TAKEOFF CONFIGURATION 

133 
IilODEL LE507A 

CONFIGURATION B tBW~~XIBp~~N~~S~~~v~~H~~G~~ 

MACH = 0.20 

RNMAC = 4.61 x lo6 

6 LE = 55Af27A 

‘FLAP = 25112.5 

iH = 0 

o( 

i 

0.01 6 - 

E 

iL 

k 

s 0.01: 2- 

!z= 

zs 

P 

i o.oot % - 

z 

0.004 

TED lb w TEU 

20 10 -10 -20 

6 

aRH 

-0.008 

0.020 - 

FIGURE 68. ROLLING-MOMENT COEFFICIENT DUE TO AILERON DEFLECTION FOR THE FCK/SLAT ,; LANDING CONFIGURATION 

134 
Spoiler effectiveness for takeoff and landing configurations is presented in 

Figures 69 and 70, respectively. The spoiler data indicated well-behaved 

characteristics for both configurations, with increasing effectiveness shown 

for increased flap deflections. The spoiler arrangement consisted of large 

chord panels compatible with space available aft of the rear spar, and 

spoiler span corresponding to flap span. This powerful spoiler 

configuration was needed because of the reduced-roll-rate capability 

associated with the high-aspect-ratio wings. 

The effect of syrunetrical spoiler deflection with landing flap deflection is 

shown in Figure 71. These results were obtained for out-of-ground-effect 

conditions. The large spoiler chord and spanwise extent was very effective 

in reducing the lift and increasing the drag; however, a significant 

positive pitching-moment shift was also apparent. Mhile the reduction in 

lift and increase in drag would result in greater deceleration on the 

ground, the positive increment of pitching moment would tend to unload the 

nose wheel. The ground effect on pitching moment, lift, and drag, with the 

spoilers deflected, should be obtained in a future test program. 

Landing Gear Effects 

The effects of the landing gear are shown in Figure 72. The gear increased 

CD by 0.0245 and decreased L/D at 1.3Vs (at CL = 1.864) from 11.55 to 

9.92. 

135 
L t ii P 4 

-I 

i 

a 

MODEL LBb07A 

CONFIGURATION B,, W,, XIB P,, N,, S,,, HID v,, 

MACH = 0.20 

6 LE 

= 8AI20A 

6 FLAP 

= 5110 

i, = O0 

0.14 

0.12 

0.08 

0.06 

0.04 

0.02 

0 

0 

0 

0 

,I 

0 

0 

0 

0 0 

0 

-0 0 

0 

ia 

0 

/’ 

‘* 

I I I I I 

0 -5 -10 -15 -20 

SPOILER DEFLECTION (DG) 

-25 -30 

FIGURE 69. ROLLING-MOMENT COEFFICIENT DUE TO SPOILER DEFLECTION FOR THE SEALED 

SLAT TAKEOFF CONFIGURATION 

136 
0.12 

0.10 

0.08 

0.06 

0.04 

0.02 

0 

MODEL LB-507A 

CONFIGURATION Bgg W lBX,B pIc N,.S,,. “ID “ICI GIA 

MACH = 0.20 

6 LE = 55A127A 

6 FLAP = 25112.5 

i, = O0 

/ 

/ 

/ 

/ 

/ 

/ 

s 

/ 

/ 0 

/ 

0 -5 -10 -15 -20 -25 -30 

SPOILER DEFLECTION (DEG) 

FIGURE 70. ROLLING-MOMENT COEFFICIENT DUE TO SPOILER DEFLECTION FOR THE FCK/SLAT 

LANDING CONFIGURATION 

137 
-. .-- 

MODEL LB-507A 

CONFIGURATION B 3BW1BX1BNlCP1CS11FGlA 

5 

w 

3. 50 

3. 00 1 

2.50- 0 

0 

0 

+ 2.00- B 0 

=I a 

tl 0 

e E I. 50- 0 

t5 o 

I. ooI3 

El 

0. !ioEl 

El 

MACH = 0.20 

RN MAC = 4.61x lo6 

6 LE = 55Al27A 

6 = 

FLAP 

25112.5 

G H 0.300- 

i 

8 

g 0.200- 

Y 

2 

El 

El 

El 

q 

El 

0 

q El 

-0. 100 El El 

q RNGLE OF$TTRK-DEG 

I3 El El 

-0.200 

0 

-0. YOO- Q o 0 0 

a 0 

0 0 

0 0 

-0. 500- 0 0 0’0 0 

0 

-0. booI es’ 

’ [ 

I I 1 I I -5 5 IO I5 20 25 0 

-0.700 

El RNGLE OF RTTRCK-DEG 

-0. 50- 

-0.800 

i 

STC 

A. LIFT AND PITCHING MOMENT 

FIGURE 71. EFFECT OF SYMMETRICAL SPOILER DEFLECTION 

i-4 

El 

El 

El 
9. ( 

j 7.c 

ie VI 

i 

. ( 

Y. 0 

3. 0 

2. 0 

) 1 

I 

MODEL LB-507A 0 0 

0 I 

a 0 

0 

0 

a 

0 

0 

0 

El 

El 

Q 0 

a 

El 

SYM RUN 6SP 

El 

I3 q 

q 

-ec! 

El a cl= 

* 0.00 

I I I I I I > I 1 I I I I I I I I 

Yt 0. OY 0. 08 0. 12 0. lb 0. 20 0. 2Lt 0. 28 0. 32 0. 37 0. YO 0. LtY 0. Y8 0. 52 0. 5b 0. b0 0. wt 0. bX 0, 

STC DRRG COEFFICIENT 

B. DRAG 

FIGURE 71. EFFECT OF SYMMETRICAL.SPOILER DEFLECTION 

72 
i 

MODEL LB-507A 

CONFIGURATION B 3BW1BX1BN1CP1CS11F 

2 

Ii 

MACH = 0.20 =1 0. 300- 

RN MAC = 4.61 x lo6 k 

3.50- 6 E 

LE = 70Ai27A 

6 5 0. 200- 

FLAP = 25112.5 

!2 

E 

0 

0 -0. IOOl 

8 

I. oo- 

-0. 2007: 

-0. 300-l 

-0. Ltoo- 

-0. 500- 

0. 50- 

I . I I I I I 

-5 0 5 IO 15 20 25 

RNGLE OF RTTRCK-DEG 

-0. 50- 

STC 

I 

RNGLE OF RTTRCK-Dz 

0 

-0. 800- 

A. LIFT AND PITCHING MOMENT 

FIGURE 72. EFFECT OF LANDING GEAR 
1 

9. 0, 

8. 0, 

7. 0 

: 

. 

; 

1 

: 

, Il.0 

3 

: 

: 5.0 

I 

v. 0 

3. 0 

2. 0 

I. 0 

--ed 

OLt c 

STC 

If-l1. c 

MODEL LB-507A El 0 

B 

El 0 

0 

I3 0 B 

0 

B 0 q 

0 0 

0 

q 

0 

I3 

0 

I3 0 SYM RUN GEAR 

I3 a , 8 :: :F;IF 

I I I I I I I I I I I I f 1 I I 1 

IO 0. OLt 0. 08 0. 12 0. lb 0. 20 0. 2* 0. 2x 0. 32 0. 3b 0. YO 0. VI 0. Ltx 0. 52 0. 5b 0. b0 0. t9t 0. bX 

DRRG COEFFICIENT 

72 

B. DRAG 

FIGURE 72. EFFECT OF LANDING GEAR 

CONCLUSIONS AND RECOMMENDATIONS 

Conclusions 

As a result of wind tunnel testing conducted at the NASA Ames 12-Foot 

Pressure Tunnel and the NASA Langley V/STOL Tunnel, the objectives set for 

the EET Phase II investigation of high-lift systems for advanced transports 

have been accomplished. This combined NASA/Douglas research effort has 

demonstrated the aerodynamic benefits of advanced-technology high-lift 

systems, has established a comprehensive data base for analysis of 

developing methods, and has identified future development areas. 

The following conclusions are drawn from the LB-486 data: 

1. Reduced VCK deflections, compared to those employed during Phase I 

testing, provided no benefit in terms of additional C LMAXor 

improved stalling characteristics. 

2. With takeoff flaps, use of a sealed outboard slat with a clean 

leading edge inboard provided significant improvement in L/D and 

pitching-moment characteristics compared to the basic slat 

configuration. This configuration resulted in a significant 

penalty in CLMAX. Use of an inboard sealed or small-gap slat at 

an intermediate deflection is a candidate for future low-speed 

testing. 

3. The full-span FCK offered no obvious advantages.in high-lift 

performance compared to either a full span VCK or a full-span slat; 

however, an FCK (especially a short-chord FCK) inboard, used in 

conjunction with a slat outboard, provided the greatest improvement 

in stalling behavior with only a relatively small loss in C LMAX. 

4. The revised slat-trim configurations tested showed less improvement 

in pitching-moment characteristics and a larger loss in C LMAX 

than the short-chord FCK/slat (inboard/outboard) combination. 

143 
5. The use of a single-segment flaperon in place of the high-speed 

aileron significantly increased C LMAX 

without penalizing L/D or 

pitching-moment characteristics. Replacement of the single-segment 

flaperon with a two-segment flaperon resulted in an additional 

small increment in maximum lift. 

6. Comparison of aerodynamic data for equivalent configurations in the 

Ames 12-Foot Pressure Tunnel and the Langley V/STOL Tunnel 

indicated generally good agreement for the lift characteristics. 

The comparisons indicated differences in pitching moment and drag. 

The following conclusions are drawn from the LB-507 data: 

1. For the high Reynolds number test condition, the cruise wing 

achieved a tail-off C 

LMAX 

of 1.59 and an L/D at 1.2V, of 

20.02. Pitch characteristics were influenced by changes in Mach 

and Reynolds number. 

2. The optimization of the leading-edge devices indicated superior 

CLMAX and pitching moments for the configurations with an inboard 

FCK; the L/D values for the inboard sealed-slat and FCK 

configurations were equivalent. The sealed-slat configurations 

exhibited lateral instability near stall under a yawed condition. 

Improvement in aerodynamic performance and pitch characteristics 

could result from further leading-edge-device optimization studies. 

3. Testing of the highly deflected flap (35"/10") indicated little 

increase in C 

LMAX' 

but a large increment in drag. 

4. Mach and Reynolds number effects were studied during the test 

program for selected configurations. CL 

MAX' 

pitching moments, 

and L/D values tended to improve with increasing Reynolds number 

and decreasing Mach number. Extrapolation of the wind tunnel data 

to flight Reynolds numbers suggested further increases in maximum 

lift are possible. 

144 
5. The nacelles and'pylons increased the cruise wing C 

the 'LMAx 

LMAX 

by 0.1; 

increment on the flaps-deflected configuration was 

nearly zero. The presence of the nacelles and pylons tended to be 

a post-stall stabilizing influence. 

6. The strakes, which were added to improve the CLmax of the 

slatted configurations, were effective in that respect. The 

additional CLmax for the inboard slat configuration with 

landing flaps was 0.14; for the takeoff flaps, 0.06. The 

strakes did not, on the other hand, increase the maximum lift 

values of the cruise wing nor of the FCK configurations. In 

all cases, the strakes were detrimental to the longitudinal 

stability. 

7. Aileron effectiveness studies indicated that, for all flap 

settings, negative deflections (trailing edge up) were more 

effective than positive deflections (trailing edge down). In 

some cases, the incremental rolling moment obtained with the 

negative aileron deflections was more than twice that obtained 

with the corresponding value for positive aileron deflection. 

8. The effect of spoiler deflection on roll characteristics 

increased as flap deflection increases. Symmetrical spoiler 

deflections for landing flap settings were very effective in 

reducing lift and increasing drag. 

Recommendations 

Analysis of the Phase II study data has identified those areas where 

continued work could result in further improvement of the technology. The 

potential for improvement has been noted in the following low-speed 

aerodynamic characteristics: pitching moments for high-lift configurations 

and increases in maximum lift for both landing and takeoff configurations. 

It is therefore recommended that future studies include the following: 

1. The use of small gaps to improve the pitching-moment 

characteristics of slat configurations without decreasing L/D. 

145 
2. The use of a slat that has a larger slot near the pylon than near 

the fuselage, to increase the section CLMAX of the inboard wing 

panel, and to promote a more rapid-inboard lift loss after stall. 

3. Additional testing of the inboard short-chord FCK, in order to 

increase the configuration L/D by reducing deflection and/or 

closing the gap. 

4. High-lift testing in ground effect at high Reynolds number. 

5. Reduced landing slat deflections to increase C 

LMAX' 

6. Higher-Reynolds-number testing to determine CLMAX and 

pitching-moment trends at conditions more closely matching those of 

flight. 

146 
B2A 

'3B 

'2B 

HIA 

"1A 

*2A 

'2A 

'1A 

GIA 

APPENDIX A 

LB-486 A,B,C 

CONFIGURATION NOTATION 

Simulates the DC-X-200 Model D-969N-21 fuselage. Full-scale 

dimensions: Length = 42.29 m (138.8 ft); constant section 

diameter = 602 cm (237 in.). The aft fuselage tail cone uses 

the DC-10 model parts. The fuselage is configured for tandem 

strut support system. 

Simulates the DC-X-200 Model D-969N-21 wing and is lofted to 

represent the airplane wing with a l-g load. Full scale 

dimensions: sW = 212.597 m2 (2288.457 ft2j; 

bW = 47.252 m (155.027 ft); aspect ratio = 10.502; 

x = 0.1407; MAC = 5.351 m (17.555 ft). The model wing has a 

removable leading edge, full-span VCK flap, trailing-edge 

two-segment flap, outboard aileron on one side, and spoilers. 

The wing is constructed of Armco 17.4 steel and contains five 

rows of pressure orifices. 

Wing-fuselage fillet for B2AW3B. 

Horizontal stabilizer for DC-X-200 (slab surface). 

Vertical stabilizer for DC-X-200 (slab surface). 

Flow-through, short core cowl nacelle configuration (2). 

New pylons for mating N2A to wing W2B (2). 

Nacelle strake configuration (attaches to N2A, 2 each 

nacelle). 

Main and nose landing gear defined for the DC-X-200'airplane. 

Main gear wheel wells with gear extended are not provided. 

147 
APPENDIX A (CONTINUED) 

a2A The outboard aileron with inboard trim at Xl{ = 89.020 cm 

(35.047 in.) and outboard trim at Xw = 109.480 cm 

(43.102 in.). The hingeline is located at 75% C. 

fl'f2 Inboard spoiler segments fabricated as individual parts. 

Superscript R = right side, L = left side, None = both sides. 

flA'f2A f, and f2 inboard O" spoilers with sheet metal aft 

extension. Trailing-edge step is filled with wax and faired 

(LB-486A). This assembly was refurbished and the T.E. step 

filled with potting (LB-486C). 

L2A 

L3A 

L4A 

Outboard spoiler segments fabricated as one piece. 

Leading-edge slat inboard of XGI = 36.367 cm (14.318 in.) and 

support at nominal gap = 2.25% C, D.H. = 2.0% C, and 

6SLAT = 25'. 

Leading-edge slat outboard of Xw = 36.367 cm (14.318 in.) 

and supported at nominal gap = 2.25% C, O.H. = 2.0% C, and 

~SLAT = 35'. 

Leading-edge variable-camber Krueger inboard of wing station 

xw 

= 36.367 cm (14.318 in.) and supported at the nominal 

gap = 2.82% C, O.H. = -0.725% C, and 6vCK = 55O. 

Leading-edge variable-camber Krueger outboard of wing station 

xw 

= 36.367 cm (14.318 in.) and supported at the nominal 

gap = 3.5% C, O.H. = 1.0% C, and 6VCK = 55O. 

L5a The inboard VCK extension to the fuselage. 

148 
APPENDI’X A (CONTINUED) 

L6A The VCK section at the pylon interruption. 

FIA Inboard main flap of a two-segment flap with inboard trim at 

xw = 13.868 cm (5.460 in.) and outboard trim at 

xW = 30.793 cm (12.123 in.). 

F2A 

F3A 

F4A 

Inboard aft flap of a two-segment flap trimned to match FIA 

and supported from FIA. 

A single-slot flaperon with inboard trim at Xw = 30.793 cm 

(12.123 in.) and outboard trim at XW = 43.411 cm 

(17.091 in.). 

Outboard main flap of a two-segment flap with inboard trim at 

xw 

= 43.411 cm (17.091 in.) and outboard trim at 

Xw = 89.020 cm (35.047 in.). 

F5A Outboard aft flap of a two-segment flap trimmed to match F4A 

and supported from FqA. 

xw ’ yw 

Wing coordinates (spanwise, chordwise). 

CIFRP Angle of attack, in degrees, of the fuselage reference plane 

relative to the equivalent free airstream. Nose up is 

positive. 

Aileron deflection, in degrees. Positive deflection is 

trailing edge down. 

GFAFT 

Aft flap deflection, in degrees (see Figure 51). 

6F MAIN 

Main flap deflection, in degrees (see Figure 51). 

'SLAT Slat deflection, in degrees (see Figure 481. 

149 
APPENDIX A (CONTINUED) 

'%CK 

iH 

sl 

s2 

s3 

s4 

s5 

VCK deflection, in degrees (see Figure 48). 

Incidence angle, in degrees, of the horizontal stabilizer 

HIA Positive deflection is trailing edge down. 

Sumnary Code 

B2AW3BX2Ba2A' Body + cruise wing. 

BWXNPZLLFFFFF 2A 3B 2B 2A 2A JA 3A 4A JA 2A 3A 4A 5A 

a2AfJ, 2, 2A fJA, 2A, 3, 4, 5, 6’ Body+fJwed wQu+VCK 

leading-edge device+flaps+nacelles, pylons, and nacelle strakes 

+VCK filler blocks. 

S2-W3B+W3D' Configuration S2 - VCK filler blocks. 

S2-W3B+W3D-fl,2 + flAfpA. Configuration 

S3+inboard spoiler trailing-edge extensions. 

B W X M P Z L L F F F F F 2A 3B 2B 2A 2A JA JA 2A JA 2A 3A 4A 5A 

a2A flA, f2A, f3, f4, f5, f6. Body+flapped wing 

+slat and WUSS leading-edge+flaps+nacelles, pylons, and nacelle 

strakes. 

150 
APPENDIX B 

LB-486A,B,C 

DIMENSIONAL DATA 

COMPONENT 

FUSELAGE (B2A) .------ 

Length 

Maximum width 

Maximum height 

WING (w3B) 

Area 

Span 

Mean aerodynamic chord 

Root chord (trapezoidal wing) 

Total root chord 

Tip chord (trapezoidal wing) 

Total tip chord 

Aspect ratio 

Taper ratio 

Spanwise station of MAC 

Fuselage station of 25% MAC 

Sweepback of 25% Cw 

Dihedral("lg") 

HORIZONTAL STABILIZER (H, A) 

Area 

Span 

MAC 

Root chord 

UNITS 

cm (in.) 

cm (in.) 

cm (in.) 

m2 (ft2) 

m (ft) 

m (ft) 

cm (in.) 

cm (in.) 

cm (in.) 

cm (in.) 

cm (in.) 

cm (in.) 

deg 

deg 

m2 (ft2) 

cm (in.) 

cm (in.) 

cm (in.) 

MODEL SCALE 

198.77 (78.255) 

28.293 (11.139) 

28.293 (11.139) 

0.4696 (5.055) 

2.221 (7.286) 

0.251 (0.825) 

37.076 (14.597) 

51.895 (20.431) 

5.217 (2.054) 

9.27 (3.65) 

10.502 

0.1407 

41.580 (16.370) 

160.28 (63.102) 

28.57 

4.5 

0.1298 (1.397) 

70.234 (27.651) 

19.91 (7.839) 

27.384 (10.781) 

151 
APPENDIX 6 (CONTINUED) 

COMPONENT 

HORIZONTAL STABILIZER (H,A) (continued) 

Tip.chord 

. Aspect ratio 

Taper ratio 

Sweepback of 25% chord : 

Dihedral 

Fuselage station of 25% HMAC 

Tail length (25% WMACto 25% HMAC) 

VERTICAL STABILIZER ($A) --... -- 

Area 

Span 

MAC 

Root. chord 

Tip chord 

-Aspect ratio 

Taper ratio 

Sweepback of 25% chord 

Tail length(25% WMAC to 25% VMAc) 

OUTBOARD AILEAR!! (azA) 

Area aft of hingeline ,, 

Span 

Chord aft of hingeline 

SPOILER (fl,f2) 

Area (each) 

Span (each) 

UNITS MODEL SCALE 

cm (in.) 9.583 (3.773) 

3.800 

0.35 

deg 30.0. 

deg 10.0 

cm (in.) 247.36 (97.384) 

cm (in.) 87.076 (34.282) 

m2 (ft2) 

cm (in.) 

cm (in.). 

cm (in.) 

cm (in.) 

deg 

cm (in.) 

0.099 -.-(1 .060) 

39.700 (15.630) 

26.731 (10.524) 

366759 (14.472) 

12.87 (5.065) 

1.6 

0.35 

35.0 

82.301 (32.402) 

cm2 (in21 

% b/2 

% C,", 

54.4 

18.4 

25.0 

(8.44) 

cm2 (in") 47.2 (7.32) 

cm (in.) 13.2 (5.18) 

152 
APPENDIX B (CONCLUDED) 

COMPONENT 

SPOILER (f3,f4,f5,f6) 

Area (total, one side) 

Span (total, one side) 

UNITS 

cm2 (in') 

cm (in.) 

NACELLE (NzA) 

Length 

Maximum cowl height 

Inlet diameter (fan cowl) 

Exit area (gas generator) 

Incidence of thrust line to FRP 

cm (in.) 

cm (in.) 

cm (in.) 

cm2 (in'.) 

deg 

Toe in deg 

MODEL SCALE 

104.660 (16.222) 

43.835 (17.258) 

32.00 (12.60) 

13.7 (5.38) 

9.85 (3.88) 

6.86 (1.06) 

1.6 

1.8 

153 

APPENDIX C 

LB-486A,B,C GRID NOTATION 

SLAT GRID NOTATION 

All gaps and overhangs are percent of local wing chord 

Dimensions are model scale 

xw 

= 14.140 cm xW = 36.367 cm 

(5.567 in.) (14.138 in.) 

'SLAT 

25' 

25O 

25O 

15O 

15O 

15O 

5O 

35O 

35O 

35O 

GAP O.H. GAP O.H. 

2.25 

3.25 

-2.0 

-1.0 

-2.0 

-2.0 

-1.0 

2.25 

1.50 1.50 

3.25 

-2.0 

-1.0 

-2.0 

2.25 2.25 

1.50 1.50 

-2.0 

-1.0 

3.25 -2.0 3.25 -2.0 

= 0.0 +7.54 = 0.0 +4.65 

2.25 2.25 

1.50 

-2.0 

-1.0 

-2.0 

1.50 

3.25 3.25 

-2.0 

-1.0 

-2.0 

155 
APPENDIX C (continued) 

LB-486A,B,C GRID NOTATION 

SLAT GRID NOTATION 

All gaps and overhangs are percent of.local,wirig-chord 

Dimensions.are,model scale 

xw =.36.367 cm xW = 89.020'cm - 

(14.138 in.) (35.047 in.) 

'SLAT - 

GAP O.H. GAP O.H. NOTATION 

25' 2.25 -2.0 2.25 -2.0 L2AD 

25' 1.50 -1.0 1.50 -1.0 L2AE 

25O 3.25 -2.0 3.25 -2.0 L2AF 

zoo -N 0 +2.0 z 0 +2.0 L?AG 

156 
APPENDIX C (continued) 

LB-486A,B,C GRID NOTATION 

VCK GRID NOTATION 

All gaps and overhangs are percent of local wing chord 

Dimensions are model scale 

xw = 14.140 cm $ = 36.367 cm 

(5.567 in.): (14.138 in.) 

'VCK GAP O.H. 'VCK - - 

GAP O.H. 

51.31a" 2.82 -0.725 55O 3.5 -1 

51.3180 2.82 -1.725 55O 3.5 -2 

51.318' 1.82 -0.725 55O 2.5 -1 

51.3180 1.82 -0.275 55O 2.5 0 

Xw = 36.367 cm xw = 111.274 cm 

(14.318 in.) (43.809 in.) 

55" 3.5 -1 55" 3.5 -2 

55O 2.5 -1 55O 2.5 0 

xW = 14.140 cm Xl4 = 36.367 cm- 

(5.567 in.) (14.318 in.) 

41.318O 2.82 -0.725 45O 3.5 -1 

41.318' 0.82 -0.725 4o" 0.5 -1 
APPENDIX C (continued) 

LB-486A,B,C GRID NOTATION 

VCK GRID NOTATION 

All gaps and overhangs are percent of local wing chord 

Dimensions are model scale 

'VCK 

41.31a" 

41.31 a0 

45O 

45O 

xW = 14.140 cm xW = 36.367 cm 

(5.567 in.) (14.318 in.1 

GAP O.H. 

- - 'VCK NOTATION 

1.82 -0.725 

1.82 -0.725 

xw = 36.367 cm 

(14.318 in.) 

3.5 -1 

2.5 -1 

45O 

45O 

45O 

45O 

2.5 -1 

2.5 0 

xW = 111.274 

(43.809 in.) 

3.5 -1 

2.5 -1 

L3AG 

L3AH 

156 
'FCK 

31.065O 

31.065' 

31.065' 

35O 

35O 

31.065O 

31.065' 

31.065' 

45O 

45O 

APPENDIX C (Continued) 

LB-486A,B,C GRID NOTATIONS 

FCK GRID NOTATIONS 

All gaps and overhangs are percent of local wing chord 

Dimensions are model scale 

xW = 14.140 cm xW = 36.368 cm 

(5.567 in.) (14.318 in.) 

GAP 

2.82 

1.82 

0.33 

2.5 

0.5 

2.82 

1.82 

0.33 

O.H. 

-0.725 

-0.725 

-0.33 

-1.0 

-0.5 

-0.725 

-0.725 

-0.33 

xW = 36.368 cm 

(14.318 in.) 

2.5 -1.0 

1.5 -1.0 

'FCK 

35O 

35O 

35O 

35O 

35O 

45O 

45O 

45O 

45O 

45O 

GAP 

3.5 

3.5 

0.5 

1.5 

0.5 

3.5 

2.5 

0.5 

O.H. 

-1.0 

-1.0 

-0.5 

-1.0 

-0.5 

-1.0 

-1.0 

-0.5 

xW = 111.036 cm 

(43.715 in.) 

2.5 -1.0 L8AD 

1.5 -1.0 L8AE 

159 
GFCK 

APPENDIX C (CONTINIIED) 

LB-486A,B,C GRID NOTATIONS 

FCK GRID NOTATIONS 

All gaps and overhangs are percent of local wing chord 

Dimensions are model scale 

xW = 36.367 cm 

(14.318 in.) 

45O 

51.065' 

51.065'. 

51.065' 

GAP O.H. 

0.5 -0.5 

2.82 -0.725 

1.82 -0.725) 

0.33 -0.33 

xW = 36.368 cm 

(14.318 in.) 

55O 2.5 -1.0 

55O 1.5 -1.0 

55O 0.5 -0.5 

xW = 14.140 cm xw = 36.368 cm 

(5.567 in.) (14.318 in.) 

5o" 

60' 

7o" 

0.05 -0.5 

0.05 -0.5 

0.05 -0.5 

XW = 111.036 cm 

(43.715 in.) 

GFCK GAP O.H. 

45O 0.5 -0.5 

55O 3.5 -1.0 

55O 2.5 -1.0 

55O 0.5 -0.5 

XW = 111.036 cm 

(43.715 in.) 

55O 2.5 -1.0 

55O 1.5 -1.0 

55O 0.5 -0.5 

NOTATION 

5o" 0.05 -0.5 L9AA 

60' 0.05 -0.5 L9AB 

7o" 0.05 -0.5 L9AC 

160 
APPENDIX C (CONTINUED) 

LB-486A,B,C GRID NOTATION 

MAIN FLAP GRID NOTATION 

All gaps and overhangs are percent of local wing chord 

Dimensions tire model scale 

Inboard Flap and Flaperon Grid 

xw .* = 14 140 cm Xw = 43.411 cm 

(5.567 in.) (17.091 in.) 

‘FMAIN 

5O 

15O 

25' 

'350 

GAP O.H. GAP O.H. 

1.3 3.2 2.5 6.0 

0.8 3.2 1.5 6.0 

0.8 2.2 1.5 4.0 

1.3 2.2 2.5 4.0 

1.6 1.1 3.0 2.0 

1.3 2.2 2.5 4.0 

0.8 2.2 1.5 4.0 

0.8 1.1 1.5 2.0 

1.6 0.0 3.0 0.0 

1.3 0.0 2.5 0.d 

1.3 0.5 2.5 1.0 

0.8 0.5 1.5 1.0 

1.9 1.1 3.5 -2.0 

1.3 0.0 2.5 0.0 

1.3 0.5 2.5 1.0 

1.i 0.5 2.0 1.0 

161 
APPENDIX C (Continued) 

LB-486A,B,C GRID NOTATIOH 

MAIN FLAP GRID NOTATIOM 

All gaps and overhangs are percent of local wing chord 

Dimensions are model scale 

OUTBOARD FLAP GRID 

‘FMAIN 

Xw = 43.411 cm 

(17.091 in.) 

xW = 89.020 cm 

(35.047 in.) 

GAP O.H. 

2.5 6.0 

5O 1.5 6.0 

1.5 4.0 

2.5 4.0 

15O 

25' 

35O 

3.0 2.0 

2.5 4.0 

1.5 4.0 

1.5 2.0 

3.0 0.0 

2.5 0.0 

2.5 1.0 

1.5 1.0 

3.5 -2.0 

2.5 0.0 

2.5 1.0 

2.0 1.0 

NOTATION 

162 
APPENDIX C (Concluded) 

LB-486A,B,C GRID NOTATION 

AFT FLAP GRID NOTATION 

All gaps and overhangs are percent of local wing chord 

Dimensions are model scale 

FLAPERON DIFFERENTIAL POSITION 

25O 

GFAFT 

xW = 14.140 cm 

(5.567 in.) 

GAP O.H. GAP O.H. NOTATION 

7.5O 0.3 0.8 0.4 1.1 

loo 0.3 0.8 0.4 1.1 

12.5' 0.4 0.4 0.5 0.5 

15O 0.4 0.4 0.5 0.5 

xW = 43.411 cm xw = 89.020 cm 

(17.091 in.) (35.047 in.) 

GFAFT GAP O.H. GAP O.H. HOTATION 

xW = 43.411 cm 

(17.091 in.) 

GAP O.H. 

2.5 1.0 

xw 

= 30.793 cm 

(12.123 in.) 

NOTATION 

F3AR 

7.5O 

loo 

12.5' 

15O 

0.5 1.5 

0.5 1.5 

0.75 0.75 

0.75 0.75 

0.5 

0.5 

0.75 

0.75 

1.5 

1.5 

0.75 

0.75 

163 

APPENDIX D 

LB-507A 

CONFIGURATION NOTATIONS 

'B3B - 

Fuselage represents the ATMR-11 aft fuselage and center body. 

The fuselage nose is the same as the one used with fuselage 

B3A* The fuselage has cutouts for the tandem-strut-support 

system and wiper for horizontal tail. Fuselage 

length = 44.2492 m (145.9619 f-t). (F.S.), constant section 

diameter = 4.310 m (14.142 ft). (F.S.). 

'1B - 

Flew technology wing, rigged to represent the airplane wing 

under a "lg" load at test conditions. Full scale trapezoidal 

dimensions: SW = 148.0 m2 ( 1600 ft'j; bW = 40.6198 m 

(133.267 ft.); AR = 11.10; x = 0.275; MAC = 4.054 m 

(13.300 ft); I? = 5O. The model has removable leading and 

trailing edges, spoilers, outboard ailerons, and four rows of 

pressure orifices. 

'1B - 

Iding fuselage fillet for B3BH1B with two strut clearance 

holes added. 

L3A - 

Inboakd conventional leading-edge slat extends from station 

X = 2.267 cm (5.758 in.) to Xw = 6.6464 cm (16.882 in.). 

The slat extends in a streamwise direction and the inboard and 

outboard trims are streamwise. The inboard slat deflections 

are 8O and 12.5' (streamwise angle). 

L4A - 

Outboard conventional leading-edge slat extends from 

xw = 6.943 cm (17.636 in.) to Xw = 17.532 cm 

(44.530 in.). The slat extends normal to the wing leading 

edge. The inboard trim is streamwise and the outboard is 

normal to the wing leading edge. The outboard slat 

deflections are 20' and 27.5' (streamwise angle). 

165 
APPENDIX D (CONTINUED) 

L5A - 

Inboard FCK with inboard trim normal to the wing leading edge 

at Xw = 2.399 cm (6.093 in.). The outboard trim is such 

that the Krueger will seal against the pylon. The inboard FCK 

deflections are 55' and 70'. 

FIA - 

Inboard main flap extends from station Xw = 1.8047 cm 

(4.584 in.) to xW = 6.883 cm (17.484 in.). The flap 

deflections are 5O, 15O, 25O, and 35'. A pressure row 

is located at Xw = 6.183 cm (15.704 in.) (left hand). 

F2A - 

Inboard auxiliary flap trim station same as FIA. The 

deflection angles of the auxiliary flap are 10' and 

12.5'. The pressure row is located at Xw = 6.183 cm 

(15.704 in.). 

F3A - 

Outboard main flap extends from station Xw = 6.895 cm 

(17.514 in.) to Xw = 14.059 cm (35.710 in.) at the flap 

leading edge and Xw = 14.133 cm (35.897 in.) at the flap 

trailing edge. The pressure rows are located at 

xW = 10.069 cm (25.575 in.) (left hand) and Xw = 12.807 cm 

(32.530 in.) (right hand). The flap deflections are 5O, 

15O, 25O, and 35'. 

F4A - 

Outboard auxiliary flap trim station XW = 6.895 cm 

(17.514 in.) to Xw = 14.133 cm (35.895 in.). 

alA - 

The trim is streamwise aft to 30% C, at which point the cut 

slants outboard to permit flap deflection. The aileron 

outboard trim station at the leading edge is 

166 
APPENDIX D (CONTINUED) 

xW = 17.661 cm (44.858 in.) and it is a streamwise cut. The 

aileron deflections available are -20°, -loo, O", 

+lO", and +2@. The aileron does not have a built in seal. 

bFIB - 

Wing flap linkage fairings representing D-3243-11 cruise 

configuration from LB-506A. Four per side in.addition to the 

fairing incorporated into pylon. 

bFIC - 

bFlB deflected to maximum position to allow flaps to 

deflect. One position only relative to the main flap. 

flA - 

One-piece bent plate representing the three inboard spoiler 

segments having a 8.66 cm (22 in.) constant chord. (F.S.). 

f2A - 

One-piece bent plate representing the three outboard segments 

having a 7.874 cm (20 in.) constant chord. (F.S.). 

GIA - 

Main and nose landing gear defined for an EET/ACA airplane. 

NIC - 

A 5.59% scale flow through nacelle representing the Pratt & 

Whitney JTlOD engine. This is the same nacelle configuration 

used with the LB-506A model. 

plc - 

A 5.59% scale pylon used in conjunction with the WIB wing 

and the NIC nacelle. The pylon positions the nacelle 

centerline at +2O with respect to the FRP and toed-in 2O 

with respect to plane of symmetry. The pylon is the same one 

used in conjunction with WTM LB-506A. 

167 
APPENDrx D (CONCLUDED) 

D2A - 

Same dorsal profile as D,B with a modified leading edge 

contour. 

HID - 

LB-506 H,C horizontal stabilizer modified at inboard end to 

match BgB fuselage. Remote control position capability. 

S = 0.1144 m* (1.2312 ft2); AR = 4.10; x = 0.350; 

sweep CV,4 = 30°; r = lO.OO. 

'1D - 

LB-506 V,c vertical stabilizer modified at the root to match 

V3B fuselage. Sv = 0.0865 m2 (0.9312 ft2); 

AR = 1.600; x = 0.35; sweep Cv,4 = 35O. 

'1lF - 

Nacelle strakes from DC-10 model LB-246 on Nlc nacelle. 

168 
APPENDIX E 

LB-507A 

DIMENSIONAL DATA 

COMPONENT 

FUSELAGE B3B 

Length 

Diameter - Constant Section 

WING W,b 

Trapezoidal gross area 

Sweepback of the quarter chord 

Taper ratio 

Aspect ratio 

Trapezoidal root chord 

Tip chord 

Mean aerodynamic 

Span 

Spanwise location of MACW 

Dihedral (lg) 

VERTICAL STABILIZER V,D 

Gross area 

Aspect ratio 

Taper ratio 

Sweepback at c/4 

Theoretical root chord 

Theoretical tip chord 

Mean aerodynamic chord 

Spanwise MACV position 

Horizontal distance 

from 25% cW to 25% c -V 

UNITS 

cm (in.) 

cm (in.) 

m2 (ft2) 

deg 

cm (in.) 

cm (in.) 

m (ft) 

m (ft) 

cm (in.) 

deg 

m2 (ft2) 

deg 

cm (in.) 

cm (in.) 

cm (in.) 

cm (in.) 

cm (in.) 

MODEL SCALE 

248.680 (97.908) 

24.079 (9.480) 

.4645 (4.9997) 

26.00 

0.275 

11.10 

32.090 (12.636) 

8.840 (3.480) 

0.2266 (0.743) 

2.271 (7.449) 

46.007 (18.113) 

5.00 

0.086 (0.931) 

1 .6 

0.35 

35.0 

34.442 (13.560) 

12.070 (4.752) 

25.054 (9.864) 

15.616 (6.148) 

100.952 (39.745) 

NOTE: All dimensions listed are in the FRP system. 

All angles listed are in the WRP system. 

169 
APPENDIX E (continued) 

COMPONENT 

HORIZONTAL STABILIZER HID 

Gross area 

Aspect ratio 

Taper ratio 

Sweepback at c/4 

Span 

Theoretical root chord 

Theoretical tip chord 

Mean aerodynamic chord 

Spanwise MACH position 

Fuselage station of (0.25)MACH 

Dihedral angle 

Horizontal distance 

from 25% cW to 25% cH 

OUTBOARD AILERON (a,A) 

Chord aft of hinge line 

Span 

INBOARD SPOILER (f,A) 

Area 

Span 

Chord 

OUTBOARD SPOILER (f2A) 

Area 

Span 

Chord 

UNITS 

m2 (ft2) 

dw 

cm (in.) 

cm (in.) 

cm (in.) 

cm (in.) 

cm (in.) 

cm (in.) 

deg 

cm (in.) 

%$, 

cm (in.) 

cm2 (in2) 

cm (in.) 

cm (in.) 

cm2 (in2) 

cm (in.) 

cm (in.) 

MODEL SCALE 

0.1144 (1.231) 

4.10 

0.35 

30.0 

68.4886 (26.964) 

24.750 (9.744) 

8.656 (3.408) 

17.983 (7.080) 

14.371 (5.658) 

243.507 (95.869) 

10.0 

124.419 (48.984) 

25.0 

22.793 (8.974) 

3.027 (1.192) 

29.538 (11.629) 

3.124 (1.230) 

3.453 (1.360) 

37.051 (14.587) 

2.841 (1.118) 

NOTE: All dimensions listed are in the FRP system. 

All angles listed are in the WRP system. 

170 
APPENDIX E (CON'TINUED) 

COMPONENT 

NACELLE (N,$ 

length 

Maximum cowl height 

Inlet diameter (fan cowl) 

Inlet area (fan cowl) 

Exit area (gas generator) 

Incidence of thrust line to FRP 

Toe in 

LEADING-EDGE SLAT (L3A, L4A) _----_--__ 

span (LsA - Inboard) 

Span (L4* - Outboard) 

Effective span 

INBOARD-MAIN FLAP (FIA) 

Area 

Span 

Root chord 

Tip chord 

Inboard trim (X,) 

Outboard trim (XW) 

INBOARD-AUXILIARY FLAP (FZA) 

Area 

Span 

Root chord 

Tip chord 

Inboard trim (X,) 

Outboard trim (X,) 

UNITS 

cm (in.) 

cm (in.) 

cm (in.) 

cm2 (in') 

cm2 (in2) 

de3 

deg 

cm (in.) 

cm (in.) 

%b/2 

cm2 (in2) 

cm (in.) 

cm (in.) 

cm (in.) 

cm (in.) 

cm (in.) 

cm2 (in2) 

cm (in.) 

cm (in.) 

cm (in.) 

cm (in.) 

cm (in.) 

MODEL SCALE 

30.526 (12.018) 

15.728 (6.192) 

9.327 (3.672) 

68.284 (10.584) 

16.258 (2.520) 

2.0 

2.0 

27.150 (10.690) 

68.199 (26.850) 

82.476 

170.291 (26.395) 

33.329 (13.122) 

5.386 (2.120) 

4.837 (1.904) 

11.643 (4.584) 

44.409 (17.484) 

105.631 (16.373) 

32.766 (12.900) 

3.225 (1.270) 

3.225 (1.270) 

11.643 (4.584) 

44.409 (17.484) 

NOTE: All dimensions listed are in the FRP system. 

All angles listed are in the WRP system. 

171 
APPENDIX E (CONCLUDED) 

COMPONENT 

OUTBOARD-MAIN FLAP (F3A) 

Area 

Span 

Root chord 

Tip chord 

Inboard trim (X.1 

Outboard trim (X,1 

UNITS MODEL SCALE 

cm2 (in21 181.997 (28.210) 

cm (in.) 46.217 (18.196) 

cm (in.) 4.831 (1.902) 

cm (in.) 3.014 (1.187) 

cm (in.) 210.168 (17.514) 

cm (in.) 90.980 (35.819) 

OUTBOARD-AUXILIARY FLAP (F4A) 

Area cm2 (in21 

Span cm (in.) 

Root chord cm (in.) 

Tip chord cm (in.) 

Inboard trim (X,1 cm (in.) 

Outboard trim (X,1 cm (in.) 

NOTE: All dimensions listed are in the FRP system. 

All angles listed are in the WRP system. 

121.052 (18.763) 

46.689 (18.382) 

3.124 (1.230) 

2.042 (0.804) 

44.485 (17.514) 

91.403 (35.985) 

172 
APPENDIX F 

GRID NOTATION LB-507A 

SLAT GRID NOTATION 

All gaps and overhang are percent of local wing chord 

Dimensions are model scale 

xw 

= 11.117 cm 

xw 

= 44.447 cm 

(4.377 in.) (17.499 in.) 

'SLAT GAP O.H. GAP O.H. 

Inboard 

8A 0.15 6.00 0.30 6.00 

8B 0.65 6.00 0.80 6.00 

12A 0.53 4.00 0.46 4.00 

12.5A 1.50 -1.00 1.50 -1.00 

Outboard 

20A 0.00 2.00 0.00 2.00 

20B 0.50 2.00 0.50 2.00 

27.5 2.25 -2.00 2.25 -2.00 

27.5B 1.5 -1 .oo 1.50 -1.00 

xw = 44.447 cm 

xw 

= 91.173 cm 

(17.499 in.) (35.895 in.) 

173 
APPENDIX F (CONTINUED) 

GRID NOTATION LB-507A 

FCK INBOARD GRID NOTATION 

All gaps and overhang are percent of local wing chord 

Dimensions are model scale 

xw = 11.117 cm xW = 44.447 cm 

(4.377 in.) (17.499 in.) 

'SLAT GAP OVERHANG 

55A 0.75 -0.75 

55B 1.50 -1.0 

70A 0.75 -0.75 

70B 1.5 -1 .o 

174 
MAIN 

'FMAIN 

APPENDIX F (CONCLUDED) 

GRID NOTATION LB-507A 

INBOARD TWO-SEGMENT FLAP (F,A/F2A) 

All gaps and overhangs are percent of local wing chord 

Dimensions are model scale 

xW = 44.447cm 

(17.449 in.) 

GAP OVERHANG 

AUX 

6FAUX - 

GAP OVERHANG 

5.0 1.50 4.0 10.0 0.50 1.50 

15.0 1.50 2.00 10.0 0.50 1.50 

25 2.50 0.00 12.5 0.75 0.75 

35.0 2.50 -1.00 12.5 0.75 0.75 

The inboard flap is rigged at the above station, and at the side of fuselage 

xw = 11.117 cm, (4.337 in.) with the same physical gap and overhang. 

The outboard flap is also rigged to the above percent gap and overhang values 

at station Xw = 44.447 cm (17.499 in.). At all stations, outboard, the gap 

and overhang are the same percentages of the local wing chord. 

175 

REFERENCES 

1. Oliver, Wayne R.: Results of Design Studies and Wind Tunnel Tests of an 

Advanced High Lift System for an Energy Efficient Transport. NASA 

CR-159389, 1980. 

2. Steckel, Doris K.; Dahlin, John A.; and Henne, Preston A.: Results of 

Design Studies and Wind Tunnel Tests of High Aspect Ratio Supercritical 

Wings for an Energy Efficient Transport. NASA CR-159332, October 1980. 

3. Crowder, J.P.: Fluorescent Mini-Tufts for Non-Intrusive Flow 

Visualization. McDonnell Douglas Report MDGJ7374, February 1977. 

4. The Staff of Douglas Aircraft Company: Selected Advanced Aerodynmamic 

and Active Control Concepts Development. Summary Report. NASA CR-3469, 

1981. 

177 
1. Rqmrt No. 2. Govanmmt Accdm No. 3. 

NASA CR-3523 

4. Tii md Subtitle 6. Raport Date 

WIND TUNNEL TESTS OF HIGH-LIFT SYSTEMS FOR ADVANCED July 1982 

TRANSPORTS USING HIGH-ASPECT-RATIO SUPERCRITICAL WINGS '* RrfwminOOr~nir8timwr 

7.kthorw John B. Allen, 8. &forming Orgsniution Report No. 

Wayne R. Oliver, and ACEE-17-FR-1608 

Lee A. Spacht 10. Wak Unit No. 

8. krfaming Organization Name and Address 

Douglas Aircraft Company 

McDonnell Douglas Corporation 11. Contract or Grant No. 

3855 Lakewood Boulevard NASl-15327 

Long Beach, CA 90846 13. Type of Report nd fkiod Covered 

12. Sponsoring Agency Name and Address 

National Aeronautics and Space Administration Contractor Report 

Washington, D.C. 20546 14. Sponsoring Agency Coda 

15. Supphnentary Notes 

Langley Technical Monitor: Thomas G. Gainer 

Final Report 

16. Abmact 

The wind tunnel testing of an advanced-technology high-lift system for a wide body 

and a narrow body transport incorporating high-aspect-ratio supercritical wings is 

described. This testing has added to the very limited low-speed high-Reynolds-number 

data base for this class of aircraft. The experimental results included the effects 

on low-speed aerodynamic characteristics of various leading- and trailing-edge 

devices, nacelles and pylons, ailerons, and spoilers, and the effects of Mach 

and Reynolds numbers. 

7. Key Words (Suggested by Author(s)) 18. Distribution Statement 

dings 

iigh aspect ratio High lift devices 

Supercritical Lift FEDD Distribution 

-ift augmentation Pitching moments 

iigh Lift Flow visualization Subject Category 02 

19. Security Qassif. lof this report) 20. Security Classif. (of this page) 21. No. of P-s 22. Rice 

Unclassified Unclassified 188 

Available: NASA’s Industrial Applications Centers NASA-Langley, 1982