Propfan test assessment testbed aircraft stability and control/performance 1/9-scale wind tunnel tests

TABLE OF C0NTEN"S Section Title Page

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iii FOREWORD ix LIST OF FIGURES 1.0 SUMMARY 2.0 INTRODUCTION 3 . 0 TEST APPARATUS 5 3.1 Wind Tunnel Models 3 . 1 . 1 High-speed Tests 3 . 1 . 2 Low-Speed Tests High-speed Flow Field Semispan Model 10 3 . 1 . 3 3 . 2 Model Propfan Blades 3.3 Isolated Propeller Model 14 3.4 Instrumentation 14 3 . 4 . 1 Model Pressure Instrumentation 14 3 . 4 . 2 Total Pressure Rakes 17 3 . 4 . 3 Propfan Hub Balances 17 3 . 4 . 4 Low-Speed, Six-Component Force Balance 19 3 . 4 . 5 High-speed, Six-Component Force Balance 19 3.4.6 Flow Survey Rakes 20 3.5 Calibrations 20 3 . 5 . 1 Propfan Airmotor Calibration 3 . 5 . 2 Flow Survey Probe Calibration 3 . 6 Wind Tunnel and Model Installations 20

3 . 6 . 1 High-speed Tests - Langley 16-Ft Transonic 20

Aerodynamics Wind Tunnel

Low-Speed Tests - Langley 4M x 7M Subsonic 2 3

3 . 6 . 2 Wind Tunnel

3 . 6 . 3 High-speed Flow Survey Tests - Lewis 23

8-Ft x 6-Ft Supersonic Wind Tunnel 4 . 0 TEST PROCEDURES 25

4.1 High-speed Tests - 16-Ft Transonic Aerodynamics 25

Wind Tunnel V

TABLE OF CONTENTS (CONTINUED) Section Title Page

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4 . 2 Low-Speed Tests - 4M x 7M Subsonic Wind Tunnel 28

4 . 3 High-speed Flow Survey Tests - 8-Ft x 6-Ft 28

Supersonic Wind Tunnel 4 . 4 Force Measurement Tares, Calibrations, and 33 Corrections 5 . 0 DATA ACCURACY 35 6 . 0 ANALYTICAL PREDICTIONS 39 7 . 0 RESULTS A N D DISCUSSIONS 41 7 . 1 Performance Data 41 7 . 1 . 1 Lift/Pitching Moment 41 7 . 1 . 2 Drag 44 7 . 2 Pressure Distributions 51 7 . 3 Stability and Control 63

7 . 3 . 1 Lift and Pitch - Low Speed 63

7 . 3 . 2

Lateral-Directional - Low Speed - Zero 65

S ides1ip

7.3.3 Lateral-Directional - Low Speed - Variable 68

Sideslip

7 . 3 . 4 Control Effectiveness - Low Speed 74

7 . 3 . 5

Effect of Nacelle Incidence - Low Speed 82

7 . 3 . 6 Lift and Pitch - High Speed 88

7 . 3 . 7

Lateral-Directional Effects - High Speed 96

7 . 3 . 8 Elevator Effectiveness - High Speed 103

7 . 3 . 9

Rudder Effectiveness - High Speed 1 0 3

7.3.10

Aileron-Spoiler Effectiveness - High 103

Speed 7 . 4 LEX (Leading Edge Extension) Performance 1 0 3 7 . 5 Wake Survey Data 107

7 . 5 . 1 Wing Pressure Measurements 107

7 . 5 . 2 Flow Field Data 1 1 1 7 . 6 Isolated Propeller Test 115 8 . 0 CONCLUDING REMARKS 127 vi TABLE OF CONTENTS (CONTINUED) T i t l e Page Sect i o n

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D E S I G N OF THE LEX (LEADING EDGE EXTENSION) 129 A P P E N D I X A APPENDIX B CALIBRATION AND DATA REDUCTION FOR 5-HOLE PROBES APPENDIX C ANALYTICAL P R E D I C T I O N S A P P E N D I X D DRAG DATA A N A L Y S I S A P P E N D I X E SYMBOLS REFERENCES vii LIST OF FIGURES Page Figure Title

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1 High Speed Model in 1 6 ' Transonic Tunnel

2 Sting Fairing - High Speed Model

3 Fuselage Cross Sections - High Speed Model

4 Low Speed Model in 4M x 7M Wind Tunnel

Balance Installation - Low Speed Model

6 High Speed Flow Field Model 7 Flow Field Model on Tunnel Side Plate Model Propfan Blade Propfan Blade Edge Thicknesses

Isolated Propeller Apparatus - High Speed Tests

11 Photograph of Isolated Propeller Model in Low Speed Wind Tunnel Model Wing Pressure Instrumentation 1 2 13 PTA Nacelle Pressure Instrumentation 14 Propfan Hub Balance 2 1 Flow Survey Rake 5-Hole Probe Effects of Transition Fix on Lift 26 Effects of Transition Fix on Drag 19 QUADPAN Panel Model of PTA Configuration 20 Comparison of QUADPAN Predictions With GI1 Data

2 1 Lift/Pitching Moment Correlation with Theory - GI1

Lift/Pitching Moment Correlation with Theory - PTA 43

23 Lift Curves f o r PTA Buildup ix LIST OF FIGURES (CONTINUED) Figure T i t l e Page

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24 PTA L i f t C o e f f i c i e n t s a t S e v e r a l Mach Numbers 46

2 5 E f f e c t of F l a p s on Lift - G I 1

E f f e c t of F l a p s on L i f t - PTA 48

27 E f f e c t of Mach Number on PTA L i f t 28 Drag P o l a r s f o r G I 1 50 5 2 29 Drag of T i p Booms 30 Power-Off Drag of PTA Nacelle 5 3 31 Drag of PTA Components a t High Speed 54 32 E f f e c t of Mach Number on PTA Drag P o l a r s 5 5 3 3 E f f e c t of Varying Nacelle Incidence on Drag 34 E f f e c t s of F l a p D e f l e c t i o n on Drag 56

3 5 Nacelle S u r f a c e P r e s s u r e D i s t r i b u t i o n s - Prop-Off 56

36 Wing S u r f a c e P r e s s u r e D i s t r i b u t i o n s - Mach 0 . 4 57

37 Wing S u r f a c e P r e s s u r e D i s t r i b u t i o n s - Mach 0 . 7

38 Nacelle S u r f a c e P r e s s u r e D i s t r i b u t i o n s - Prop-on, 59

Mach 0 . 7

Nacelle S u r f a c e P r e s s u r e D i s t r i b u t i o n s - Prop-on, 60

Mach 0 . 8

E f f e c t of PTA Nacelle on Wing S e c t i o n P r e s s u r e s - 60

Mach 0 . 7

41 E f f e c t of PTA Nacelle on Wing S e c t i o n P r e s s u r e s - 61

Mach 0 . 8 42 E f f e c t of Angle of Attack on Wing S e c t i o n P r e s s u r e s 6 2 43 E f f e c t of Mach Number on Wing S e c t i o n P r e s s u r e s 62 44 E f f e c t s of PTA M o d i f i c a t i o n s on L i f t and P i t c h i n g 64 Moment X LIST OF FIGURES (CONTINUED) Title Page Figure

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45 Effects of Propfan Power on Aerodynamic

Characteristics in Pitch - Flaps Up

46 Effects of PTA Modifications on Side Force, Yawing 67

Moment, and Rolling Moment - Flaps Up

47 Effects of PTA Modifications on Side Force, Yawing

Moment, and Rolling Moment - Flaps 20-Degrees

Effects of PTA Modifications on Side Force, Yawing 70

Moment, and Rolling Moment - Flaps 40-Degrees

Effects of Power on Aerodynamic Characteristics in 71

Pitch - Flaps Up

50 Effects of PTA Modifications on Sideslip

Characteristics - Flaps Up

51 Effects of Propfan Power on Aerodynamic

Characteristics in Sideslip - Flaps Up

52 Effects of PTA Modifications on Sideslip

Characteristics - Flaps 20-Degrees

Effects of PTA Modifications on Sideslip 53

Characteristics - Flaps 40-Degrees

Elevator Effectiveness

78 Rudder Effectiveness in Pitch - Flaps Up

56 Rudder Effectiveness in Pitch - Flaps 20-Degrees

80 Rudder Effectiveness in Pitch - Flaps 40-Degrees

8 1

58 Spoiler Effectiveness in Pitch - Flaps Up

83 Rolling Moment from Aileron and Spoiler

Deflections - Flaps Up

60 Spoiler Effectiveness in Pitch - Flaps 20-Degrees

61 Spoiler Effectiveness in Pitch - Flaps 40-Degrees

86 Effects of Nacelle Incidence on Lift and Pitching

Moment - Flaps Up

xi - & L f x b m w LIST OF FIGURES (COIJTIIWED) Figure Title Page

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63 Effects of Nacelle Incidence on Lift and Pitching 87

Moment - Flaps Up, Propfan Power On

64 PTA Configuration Buildup - Mach 0 . 4 89

65 PTA Configuration Buildup - Mach 0 . 7 90

PTA Configuration Buildup - Mach 0 . 8

66 91

PTA Configuration Buildup - Mach 0.85

67 92 Mach Number Effects on Lift Characteristics 68 93 69 Comparison of Lift Data from Low- and High-speed 94 Tests 70 Comparison of Pitch Data from Low- and High-speed 95 Tests

Comparison of GI1 and PTA in Sideslip - Mach 0 . 4

71 97

72 Comparison of GI1 and PTA in Sideslip - Mach 0 . 7

Comparison of GI1 and PTA in Sideslip - Mach 0.8

73 99

74 Comparison of G I 1 and PTA In Sideslip - Mach 0.85 100

75 Effect of Mach Number on Sideslip Derivatives 101 76 Effect of Mach Number on Lateral-Directional Offsets 102 77 Comparison of Measured and Predicted Rudder 104 Effectiveness Aileron-Spoiler Effectiveness 104 Effect of LEX on Wing Surface Pressures, Mach 0 . 7 0 105 Effect of LEX on Wing Surface Pressures, Mach 0.80 106 81 Effect of LEX on PTA Drag Pressure Distributions Inboard of Nacelle 109 Pressure Distributions Outboard of Nacelle 110 84 Rake Data for Axial Velocity Component U at Mach 0 . 6 112 x i i LIST OF FIGURES (CONTINUED) Figure Title Page 85 Rake Data for Lateral Velocity Component V at 113 Mach 0 . 6 86 Rake Data for Vertical Velocity Component W at 114 Mach 0.6 87 Rake Data for Axial and Lateral Velocity Components at Mach 0.85 88 Rake Data for Vertical Velocity Component W at Mach 0.85 89 Thrust Coefficient Data at Mach 0.4 90 Power Coefficient Data at Mach 0.4 91 Thrust Coefficient Data at Mach 0.165 92 Power Coefficient Data at Mach 0.165 93 Propeller at Angle of Attack 94 Changes in Blade Angle of Attack Due to Thrust Axis 123 Inclination 95 Cyclic Variation of Blade Environment Due to Thrust Axis Inclination 96 Thrust Coefficient Versus Angle of Attack 97 Thrust Coefficient Versus Advance Ratio at Two Angles of At tack 98 Collapse of Thrust Coefficient Data on J Cos 99 Normal Force Coefficient Data Versus Angle of Attack 125 xiii 1 . 0 SUMMARY Full-span wind t u n n e l model tests a t 1/9-scale were performed t o air- e s t a b l i s h s a f e t y of f l i g h t and performance p r e d i c t i o n s f o r t h e PTA c r a f t . Semispan model t e s t s a t t h e same scale were performed t o measure f l o w f i e l d p r o p e r t i e s ( l o c a l v e l o c i t y and f l o w a n g u l a r i t y ) i n t h e r e g i o n s where t h e propfan o p e r a t e s , To understand and be a b l e t o scale d a t a from t h e powered model tests, t h e 1/9-scale SR7L propfan r o t o r was a l s o t e s t e d on an i s o l a t e d n a c e l l e .

The s a f e t y of f l i g h t and performance d a t a were t a k e n i n t h e NASA- Langley 4 M x 7M Subsonic Wind Tunnel a t Mach 0.167 and 0.2 and i n t h e NASA-Langley 16-Ft T r a n s o n i c Aerodynamics Wind Tunnel a t Mach 0.4 t o 0.85.

The flow survey d a t a were t a k e n i n t h e NASA-Lewis 8-Ft x 6-Ft Supersonic Wind Tunnel a t Mach 0.6 t o 0.85. Reynolds numbers based on MAC ranged from approximately 0 . 2 x 106 t o 5 x 106.

The models were f u l l - s p a n and sting-mounted f o r a l l tests e x c e p t t h e high-speed flow survey tests. For t h e l a t t e r , t h e model was modified t o a h a l f - f u s e l a g e , semispan c o n f i g u r a t i o n and mounted from t h e t u n n e l s i d e w a l l .

Major emphases i n t h e f l i g h t s a f e t y l p e r f o r m a n c e tests were on wing s u r f a c e p r e s s u r e d i s t r i b u t i o n s and aerodynamic f o r c e s and moments. The g e n e r a l philosophy of t h e s e tests was t o o b t a i n d a t a from the G I 1 a i r p l a n e model f o r a b a s e l i n e and e s t a b l i s h increments from t h a t b a s e l i n e a s t h e PTA i n s t a l l a t i o n was b u i l t up. These increments were t h e n a p p l i e d t o G I 1 f l i g h t test d a t a t o p r e d i c t PTA performance.

e s t a b l i s h e d As expected, t h e a d d i t i o n of t h e PTA hardware t o t h e G I 1 a i r c r a f t : o I n c r e a s e d d r a g s i g n i f i c a n t l y o I n c r e a s e d wing l i f t and nose-up p i t c h i n g moment ( t h e r e b y s l i g h t l y reducing p i t c h s t a b i l i t y o I n c r e a s e d side f o r c e due t o s i d e s l i p o Decreased yawing and r o l l i n g moments due t o s i d e s l i p o S l i g h t l y a f f e c t e d rudder and e l e v a t o r e f f e c t i v e n e s s o S i g n i f i c a n t l y reduced r o l l c o n t r o l power o Reduced maximum l i f t c o e f f i c i e n t s G e n e r a l l y , wind t u n n e l d a t a were p r e d i c t e d w i t h good accuracy by t h e i n v i s c i d panel code QUADPAN. T h i s was t r u e n o t o n l y of model f o r c e s and moments and s u r f a c e p r e s s u r e d i s t r i b u t i o n s b u t a l s o of t h e flow f i e l d p r o p e r t i e s o f f t h e s u r f a c e i n t h e region of t h e propfan.

These model tests, t o g e t h e r w i t h supplementary QUADPAN a n a l y s e s , provided t h e d a t a t o v e r i f y from s t a b i l i t y , c o n t r o l , h a n d l i n g , and per- formance s t a n d p o i n t s t h a t t h e proposed PTA a i r c r a f t would be a s u i t a b l e v e h i c l e f o r f l i g h t r e s e a r c h tests of t h e l a r g e - s c a l e SR-7L propfan.

2.0 INTRODUCTION The NASA-Lewis Propfan T e s t Assessment (PTA) Program was i n i t i a t e d i n 1984 t o develop and f l y a t e s t b e d a i r c r a f t f o r e v a l u a t i o n of t h e Large- S c a l e Advanced P r o p e l l e r (LAP) a l r e a d y under development a t Hamilton Standard. A c o n t r a c t was awarded t o t h e Lockheed-Georgia Company f o r t h e d e s i g n and m o d i f i c a t i o n r e q u i r e d t o c o n v e r t a Gulfstream Aerospace Corporation G I 1 a i r c r a f t t o t h e PTA c o n f i g u r a t i o n . The PTA c o n f i g u r a t i o n proposed c o n s i s t e d of i n s t a l l a t i o n of an engine n a c e l l e on t h e left-hand wing of t h e GII, a d d i t i o n of an a c o u s t i c i n s t r u m e n t a t i o n boom outboard of t h e n a c e l l e , and a d d i t i o n of wing t i p booms f o r s t a t i c and dynamic balance. These m o d i f i c a t i o n s i n t u r n r e q u i r e d c o n f i g u r a t i o n changes such as d e a c t i v a t i o n of some wing s p o i l e r panels.

To a s s u r e s a f e t y of f l i g h t and v a l i d a t e performance p r e d i c t i o n s of t h e PTA a i r c r a f t , small-scale wind t u n n e l tests were planned. Tests were needed a t low-speed f l i g h t c o n d i t i o n s and a t speeds up t o Mach 0.85.

Force and moment d a t a were needed f o r performance and handling q u a l i t i e s e v a l u a t i o n , and s u r f a c e s t a t i c p r e s s u r e d a t a were needed t o e x p l a i n flow phenomena and v a l i d a t e l o a d s p r e d i c t i o n s .

A major o b j e c t i v e of t h e PTA Program is t o measure v i b r a t o r y stresses on t h e LAP r o t o r b l a d e s i n f l i g h t and compare them w i t h p r e d i c t e d values.

S i n c e t h e f i r s t s t e p i n t h e blade stress p r e d i c t i o n p r o c e s s is t o p r e d i c t t h e blade flow environment, a f u r t h e r o b j e c t i v e of t h e small-scale tests w a s t o measure flow f i e l d d a t a i n t h e v i c i n i t y of t h e propfan r o t o r and compare t h o s e d a t a w i t h p r e d i c t e d values.

A new 1/9-scale model w a s b u i l t f o r t h e s e tests. Maximum scale was l i m i t e d by t h e s i z e s of a v a i l a b l e wind t u n n e l s and minimum scale by t h e d i f f i c u l t i e s of modeling small propfan r o t o r s and d r i v e systems. The 1/9- scale set t h e r o t o r diameter a t one f o o t and r e s u l t e d i n a model small enough t o f i t i n t o s e v e r a l s u i t a b l e wind tunnels. For a l l e x c e p t t h e flow survey tests, t h e model was s t i n g mounted w i t h compressed a i r brought aboard t h e model through t h e s t i n g t o power t h e a i r motor t h a t drove t h e propfan. A v a i l a b i l i t y of d r i v e motors made i t d e s i r a b l e t o test t h e model w i t h propfan on t h e r i g h t hand wing i n s t e a d of t h e l e f t hand a s i n f u l l scale. T h i s , however, creates no problems s i n c e a m i r r o r image configura- t i o n of t h e f u l l - s c a l e a i r c r a f t was always used i n t h e scale-model tests.

Tunnel a v a i l a b i l i t y t i m e d i d n o t p e r m i t o b t a i n i n g flow f i e l d d a t a i n t h e f u l l - s p a n , high-speed wind t u n n e l tests. T h e r e f o r e , a second series of high-speed tests were performed i n a smaller wind t u n n e l s p e c i f i c a l l y t o o b t a i n flow f i e l d data. For t h e s e tests, a new h a l f - f u s e l a g e was b u i l t , and a semispan model was assembled from p a r t s of t h e f u l l - s p a n model.

The high-speed wind t u n n e l tests were conducted i n t h e NASA-Langley 16-Ft Transonic Aerodynamics Wind Tunnel i n t h e summer and f a l l of 1985.

Low-speed tests were performed i n t h e NASA-Langley 4 M x 7 M Subsonic Wind Tunnel i n t h e summer of 1986. The high-speed flow s u r v e y tests were performed i n t h e NASA-Lewis 8-Ft x 6-Ft Supersonic Wind Tunnel i n January 1987. Test Mach numbers ranged from 0.16 t o 0.85. Reynolds numbers,

based on MAC, ranged from 0.2 x lo6 i n t h e low-speed wind t u n n e l t o

5 x lo6 i n t h e high-speed wind tunnels.

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A l l of t h e d a t a c o l l e c t e d i n t h e s e tests have been s t o r e d i n informal r e p o r t s a t t h e Lockheed Aeronautical Systems

Company - Georgia as p a r t of

t h e PTA documentation.. This r e p o r t p r e s e n t s a n a l y s e s of t h o s e d a t a and t h e major r e s u l t s from t h o s e tests. Where a p p r o p r i a t e , t h e d a t a are compared with p r e d i c t e d v a l u e s from a n a l y t i c a l I n t h e case of t h e codes.

flow f i e l d d a t a , a major emphasis is placed on t h e c o r r e l a t i o n of experimental d a t a w i t h a n a l y s e s and m o d i f i c a t i o n of t h e a n a l y t i c a l methods needed t o produce good c o r r e l a t i o n .

3.0 TEST APPARATUS 3.1 WIND TUNNEL MODELS 3.1.1 High-speed Tests A photograph of t h e sting-mounted, high-speed model assembly, as i n s t a l l e d i n t h e NASA-Langley 16-Ft Transonic Aerodynamics Wind Tunnel, i s shown i n Figure 1.

A c y l i n d r i c a l f a i r i n g was a t t a c h e d t o t h e a f t p o r t i o n of t h e model as shown i n Figure 2 t o a c c e p t t h e support s t i n g . T h i s s t i n g f a i r i n g method and was recommended by f a c i l i t y person- i s commonly used i n t h i s f a c i l i t y n e l as t h e b e s t method of minimizing s t i n g attachment c o r r e c t i o n s .

The model w a s designed t o measure propfan a c o u s t i c s d a t a as w e l l as aerodynamics data. One wing contained a n a r r a y of a c o u s t i c s t r a n s d u c e r s and t h e o t h e r an a r r a y of p r e s s u r e o r i f i c e s . The propfan d r i v e motor t h a t w a s a v a i l a b l e a t t h e s t a r t of t h e program and t h e matrix of propfan r o t a - t i o n d i r e c t i o n s d e s i r e d i n t h e tests made it expedient t o d e d i c a t e t h e left-hand wing t o a c o u s t i c s i n s t r u m e n t a t i o n and t h e right-hand wing t o p r e s s u r e i n s t r u m e n t a t i o n . The aerodynamic d a t a , t h e r e f o r e , are f o r a c o n f i g u r a t i o n t h a t i s a m i r r o r image of t h e PTA c o n f i g u r a t i o n - - t h a t is, t h e n a c e l l e i s mounted on t h e right-hand wing i n s t e a d of t h e l e f t . I n t h e tests, t h e t i p booms (which are d i f f e r e n t f o r l e f t and r i g h t wings) and t h e a c o u s t i c boom were always arrayed t o maintain t h e m i r r o r image config- u r a t i o n .

Dimensions and contours of t h e b a s i c GI1 c o n f i g u r a t i o n were o b t a i n e d from l o f t drawings provided by Gulfstream Aerospace. The 1/9-scale model t h u s d e r i v e d had a wing span of 2.337m (92 i n c h e s ) and l e n g t h ( t o t h e end of t h e s t i n g f a i r i n g ) of 2.354m (92.67 i n c h e s ) . Model weight was approx- i m a t e l y 544 kg (1200 pounds). Wing area was 0.918 m 2 (9.877 f t z ) , and t h e MAC was 0.416m (16.364 inches).

Figure 3 p r e s e n t s c r o s s s e c t i o n s through t h e f u s e l a g e t o show t h e balance, compressed a i r d u c t s f o r t h e propfan d r i v e motor, scani-valve l o c a t i o n s , and o t h e r d e t a i l s of model c o n s t r u c t i o n . The balance, of c o u r s e , s e p a r a t e d t h e metric and non-metric parts of t h e model assembly.

A i r f o r t h e d r i v e motors flowed through t h e c e n t e r of t h e s t i n g , e x i t e d t h e s t i n g through p o r t s on e i t h e r s i d e of t h e a d a p t o r a t p o i n t (11, flowed through bypass l i n e s around t h e balance ( S e c t i o n B ) , and i n t o t h e forward plenum a t p o i n t ( 2 ) ( S e c t i o n A). From t h e forward plenum, t h e a i r supply could be ported t o e i t h e r / o r both wing r o o t r e g i o n s , t h e n through passages i n t h e wing t o t h e propfan n a c e l l e s .

Scani-valves f o r p r e s s u r e measurements were l o c a t e d i n t h e nose por- t i o n of t h e model. Power and s i g n a l l e a d s from t h e scani-valves passed l o o s e l y around t h e balance and were taped t o t h e e x t e r n a l s u r f a c e of t h e s t i n g t o e x i t t h e test s e c t i o n .

The right-hand wing w a s f i t t e d w i t h an a i l e r o n t h a t could be set a t z e r o and f 1 0 d e g r e e s d e f l e c t i o n by means of f i x e d b r a c k e t s . An outboard s p o i l e r ( t h e only one a c t i v e on t h e PTA a i r c r a f t ) was provided w i t h d e f l e c t i o n a n g l e s of 10 and 35 degrees.

The PTA n a c e l l e w a s c o n s t r u c t e d s o t h a t n a c e l l e tilt a n g l e s of +2, -1, and -3 d e g r e e s could be -simulated. On t h e PTA a i r p l a n e t h e n a c e l l e F i g u r e 1 . High S p e e d Model in 16' T r a n s o n i c Tunnel F i g u r e 2 . S t i n g Fairing - High Speed BALANCE ,SUPPORT FUSELAFE SHELL A-A D-D B-B c-c Figure 3 . Fuselage Cross Sections - High Speed Model breaks a t t h e i n t e r f a c e between t h e Quick-Engine-Change (QEC) assembly and t h e a f t n a c e l l e s o t h a t o n l y t h e QEC is t i l t e d . T h i s w a s t o o d i f f i c u l t t o model, however, because of t h e requirement f o r compressed air t o a plenum i n t h e model n a c e l l e t h a t extends f o r e and a f t of t h e QEC break.

T h e r e f o r e , on t h e model t h e e n t i r e n a c e l l e was t i l t e d by p u t t i n g wedges between t h e wing and the nacelle. O n t h e model, t h e f u l l - s c a l e a f t n a c e l l e was simulated o n l y for n a c e l l e tilt of +2 degrees and w a s r a i s e d above t h e t r u e f u l l - s c a l e p o s i t i o n f o r -1 d e g r e e and -3 degrees. The n a c e l l e p i v o t p o i n t on t h e model, however, w a s t h e same as for t h e air- plane s o t h a t t h e forward n a c e l l e (and propfan p o s i t i o n ) was properly simulated at a l l tilt angles.

The PTA n a c e l l e c o n t a i n e d a f o r c e balance between t h e d r i v e motor and t h e p r o p e l l e r hub s o t h a t f o r c e s and moments on t h e p r o p e l l e r could be - measured.

The h o r i z o n t a l s t a b i l i z e r was made so t h a t it could be p o s i t i o n e d a t i n c i d e n c e a n g l e s of 0, -1.5, -3, and -5 degrees. I n a d d i t i o n , e l e v a t o r of +5, +lo, -15, -20, and -25 degrees could be s e t . The d e f l e c t i o n a n g l e s rudder was made s o t h a t it could be s e t a t d e f l e c t i o n of * 5 , f 1 0 , *15, *20, and i 2 5 degrees.

Throughout t h e d e s i g n phase of t h e PTA Program, t h e r e was concern t h a t unforeseen and u n p r e d i c t a b l e t r a n s o n i c d r a g a s s o c i a t e d w i t h t h e PTA n a c e l l e i n s t a l l a t i o n might prevent a t t a i n m e n t of t h e d e s i r e d high-speed study was performed t o t e s t c o n d i t i o n s . To i n s u r e a g a i n s t t h i s , a d e s i g n d e v i s e c o n f i g u r a t i o n changes t h a t could be t e s t e d i n t h e wind tunnel and promising m o d i f i c a t i o n then a p p l i e d t o t h e a i r c r a f t i f needed. The most f o r d e a l i n g w i t h e x c e s s i v e d r a g emerged as a wing l e a d i n g edge e x t e n s i o n (LEX) on t h e inboard s i d e of t h e n a c e l l e . This d e v i c e was designed t o extend t h e l e a d i n g edge and camber it s o t h a t t h e upwash induced by pro- p e l l e r s l i p s t r e a m s w i r l would not r e s u l t i n e x c e s s i v e v e l o c i t y peaks on A more d e t a i l e d d i s c u s s i o n of t h e LEX design i s presented i n t h e wing.

Appendix A.

3.1.2 Low-Speed T e s t s The model and its i n s t a l l a t i o n i n t h e 4M x 7M Subsonic Wind Tunnel 4 . The s t i n g support and balance system are d i f f e r e n t a r e shown i n Figure from t h o s e used i n t h e high-speed tests because t h e s e items a r e g e n e r a l l y unique t o t h e test f a c i l i t y . I n t h e low-speed t e s t s , t h e s t i n g contained a tube, w i t h bellows a t e i t h e r end, t h a t c a r r i e d compressed a i r t o t h e A s shown i n Figure 5, t h e a i r passage l e d s t r a i g h t i n t o t h e model, model.

was o f f s e t from t h e s t i n g c e n t e r l i n e . This arrangement and t h e balance r e q u i r e d d i f f e r e n t a d a p t e r s and s t i n g attachment f i t t i n g s i n t h e model from t h o s e used i n t h e high-speed t e s t s .

The only o t h e r d i f f e r e n c e between high-speed and low-speed models was t h a t t h e low-speed model contained p r o v i s i o n s f o r wing f l a p extensions.

Flaps were made for each wing w i t h b r a c k e t s f o r f l a p d e f l e c t i o n s of 10 d e g r e e s , 20 d e g r e e s , and 40 degrees. It w a s necessary t o disconnect were p r e s s u r e o r i f i c e t u b e s t o t h e a i l e r o n and s p o i l e r when t h e f l a p s d e f l e c t e d .

F i g u r e 4. Low Speed Model in 4M x 7M Wind T u n n e l

t 7 1

7-1 I S U P P O R T BELLOWS / A I R FLOW \ A T T ~ C I I E S BALANCE T O hlOUCL F i g u r e 5. Balance i n s t a l l a t i o n - Low Speed Model 3.1.3 High-speed Flow F i e l d Sernispan Model The model f o r t h e high-speed flow f i e l d t e s t s ( F i g u r e 6 ) i n t h e NASA- L e w i s 8-Ft x 6-Ft Supersonic Wind Tunnel was a half-model v e r s i o n of t h e f u l l - s p a n model used i n t h e aerodynamic tests. A new h a l f - f u s e l a g e w a s b u i l t w i t h p r o v i s i o n s f o r attachment t o a model-centerline r e f l e c t i o n was mounted on one tunnel s i d e w a l l . Since t h e p l a n e , which, i n t u r n , model wing was b u i l t i n s e p a r a t e left-hand and right-hand p a r t s , it w a s a simple matter t o b u i l d an a d a p t o r t o mount o n l y t h e right-hand wing t o t h e h a l f - f u s e l a g e and s l d e w a l l p l a t e . The Spey e n g i n e n a c e l l e was r e t a i n e d as a n e s s e n t i a l p a r t of t h i s model, b u t v e r t i c a l and h o r i z o n t a l t a i l compo- n e n t s were omitted. The l a t t e r were judged t o have l i t t l e i n f l u e n c e on t h e f l o w i n t h e region around t h e f r o n t of t h e PTA n a c e l l e .

The PTA n a c e l l e and o t h e r right-hand wing accouterments were a t t a c h e d as i n t h e f u l l - s p a n model. Since t h e o b j e c t i v e s of t h e test were t o o b t a i n flow f i e l d d a t a w i t h o u t t h e propfan, no p r o v i s i o n s were made t o b r i n g compressed a i r aboard t h e model.

The s i d e w a l l mounting p l a t e (provided by t h e test f a c i l i t y ) was designed t o f i t t h e l e f t hand t u n n e l w a l l (as viewed from upstream). To test t h e right-hand h a l f of t h e model, i t was n e c e s s a r y t o mount t h e model was a t t a c h e d t o a t u r n t a b l e imbedded i n t h e s i d e w a l l i n v e r t e d . The model (as shown i n Figure 7 ) t h a t could be r o t a t e d t o change model angle p l a t e was r o t a t e d about t h e - 2 5 MAC point.

of a t t a c k . The model 3.2 MODEL PROPFAN BLADES The o b j e c t i v e i n d e s i g n of t h e model propfan b l a d e s w a s t o produce small-scale r o t o r s t h a t would perform aerodynamically l i k e t h e f u l l - s c a l e Hamilton Standard SR-7L propfan. The f u l l - s c a l e b l a d e s are f l e x i b l e enough s o t h a t they t w i s t and bend under f l i g h t loads. The small-scale b l a d e s , on t h e o t h e r hand, are r e l a t i v e l y s t i f f . The d e s i g n procedure chosen was t o make t h e small-scale b l a d e s i n t h e shape of t h e loaded f u l l - scale b l a d e i n t h e d e s i g n c r u i s e condition. Coordinates of t h e f u l l - s c a l e b l a d e s i n t h i s f l i g h t c o n d i t i o n were reduced t o model scale by Hamilton Standard and provided f o r b l a d e design.

The small s i z e of t h e model b l a d e s r e s u l t e d i n a s i t u a t i o n a t t h e b l a d e l e a d i n g and t r a i l i n g edges where d e p a r t u r e s from scale were neces- s i t a t e d by t h e f a b r i c a t i o n technique. Metal b l a d e s were not allowed i n t h e 16-Ft Transonic Aerodynamics Wind Tunnel f o r f e a r of t h e damage p o t e n t i a l i f a blade were l o s t a t t h e h i g h r o t a t i o n a l speeds ( t o about 19,000 rpm) d u r i n g t h e high-speed tests . The b l a d e s were f a b r i c a t e d , t h e r e f o r e , of carbon f i b e r composite material.

The b e s t a v a i l a b l e material was graphite-epoxy t a p e i n 0.0076 c m (0.003-inch) t h i c k n e s s . S t r e n g t h c o n s i d e r a t i o n s r e q u i r e d t h a t t h e model b l a d e s i n t h e i r t h i n n e s t r e g i o n s c o n t a i n a t l e a s t two t a p e l a y e r s w i t h t h e p l i e s o r i e n t e d a t some a n g l e r e l a t i v e t o one another. T h e r e f o r e , t h e minimum blade element t h i c k n e s s was of t h e o r d e r of 0.0152 cm (0.006 i n c h ) . A t y p i c a l b l a d e is shown i n Figure 8.

F i g u r e 9 shows l e a d i n g edge r a d i i and t r a i l i n g edge t h i c k n e s s of t h e LAP b l a d e s reduced t o 1/9-scale. It can be s e e n t h a t f o r a minimum i n c h ) , t h e l e a d i n g edge r a d i i goes material t h i c k n e s s of 0.0152 c m (0.006 o u t of s c a l e outboard of t h e 80-percent span s t a t i o n , and t h e t r a i l i n g Figure 6 . High Speed Flow Field FLO-THRU C L Figure 7 . Flow Field Model on Tunnel Side Plate F i g u r e 8 Model Propfan Blade THICK'iESS A T 5 0 % C H O R D .0.5

-\I

\ 31' 0 . 4 0 . 6 0.3 1 .a 0 . 1 F R A C T I O N OF T I P R A D I U S F i g u r e 9 . Profan Blade Edge T h i c k n e s s e s edge t h i c k n e s s i s o u t of scale over t h e whole b l a d e span. It was expected t h a t the e x c e s s l e a d i n g edge t h i c k n e s s would be more c r i t i c a l than t h e excess t r a i l i n g edge t h i c k n e s s and t h a t t h e major impact would be on r o t o r e f f i c i e n c y a t high speeds. Because t h i s i s t h e speed r e g i o n where propfan power e f f e c t s on t h e a i r c r a f t are l i k e l y t o be smallest, i t w a s b e l i e v e d t h a t t h e out-of-scale t h i c k n e s s could be t o l e r a t e d .

During t h e c o u r s e of t h e high-speed tests, t h e r e were two r o t o r f a i l u r e s , T h i s r e s u l t e d u l t i m a t e l y i n a r e d e s i g n of t h e b l a d e s t o t h i c k e n t h e region n e a r t h e shank where f a i l u r e s were b e l i e v e d t o i n i t i a t e . These b l a d e s were used i n t h e low-speed wind t u n n e l tests, and t h e r e were no more b l a d e f a i l u r e s .

The b e s t measure of how w e l l t h e b l a d e s were s c a l e d d e r i v e s from r e s u l t s of t h e i s o l a t e d p r o p e l l e r tests. It w i l l be shown i n d i s c u s s i o n of t h o s e r e s u l t s t h a t t h e model propfan r o t o r s simulated t h e f u l l - s c a l e a r t i c l e reasonably w e l l .

3 . 3 ISOLATED PROPELLER MODEL The o b j e c t i v e i n d e s i g n i n g a p p a r a t u s f o r i s o l a t e d p r o p e l l e r tests w a s t o produce a test n a c e l l e and support t h a t i n t e r f e r e s minimally w i t h t h e p r o p e l l e r . Apparatus f o r t h e high-speed i s o l a t e d n a c e l l e tests is shown i n F i g u r e 10. The hub, s p i n n e r , b a l a n c e , and motor were t h e same u n i t s used on t h e a i r c r a f t model i n s t a l l a t i o n . The n a c e l l e , however, was a s t r e a m l i n e d axisymmetric c y l i n d e r supported by an aerodynamically con- t o u r e d , s w e p t s t r u t connected t o t h e model s u p p o r t s t i n g . Drive motor a i r was r o u t e d through t h e s t i n g and support s t r u t t o t h e forward p a r t of t h e n a c e l l e , through t h e motor, and t h e n o u t t h e a f t p o r t i o n of t h e n a c e l l e .

To measure t h e aerodynamic f o r c e s on t h e p r o p e l l e r , t h e n a c e l l e was f i t t e d w i t h t h e same hub balance used i n t h e a i r c r a f t model--located between t h e p r o p e l l e r hub and t h e d r i v e motor.

Other i n s t r u m e n t a t i o n included s t a t i c p r e s s u r e p o r t s upstream and downstream of t h e motor t o measure s t a t i c p r e s s u r e d r o p a c r o s s t h e motor, a t o t a l p r e s s u r e rake a t t h e n a c e l l e e x i t , and thermocouples a t t h e motor i n l e t , motor o u t l e t , and downstream of t h e choke p l a t e .

The a p p a r a t u s used f o r t h e low-speed i s o l a t e d p r o p e l l e r tests i s shown i n F i g u r e 11. Design requirements f o r t h e low-speed t u n n e l d i c t a t e d a s t i f f e r n a c e l l e s u p p o r t than t h a t used i n the high-speed tests. Thus, t h e c o n f i g u r a t i o n shown i n F i g u r e 11 used two c y l i n d r i c a l h a l v e s t o clamp t h e n a c e l l e . The lower h a l f was welded t o an aerodynamically contoured v e r t i c a l s t r u t w i t h a bottom p l a t e t h a t b o l t e d d i r e c t l y t o t h e NASA model s u p p o r t f i x t u r e . The b u l k i n e s s of t h e i s o l a t e d n a c e l l e model f o r t h e s e low-speed tests was g r e a t e r t h a n d e s i r e d , b u t was deemed a c c e p t a b l e because i t was a low speed test.

I n s t r u m e n t a t i o n was the same as for t h e high-speed i n s t a l l a t i o n .

3.4 INSTRUMENTATION 3.4.1 Model Pressure I n s t r u m e n t a t i o n The right-hand wing of t h e model was instrumented w i t h 196 s t a t i c p r e s s u r e o r i f i c e s as shown i n F i g u r e 12. These were l o c a t e d a t spanwise s t a t i o n s such t h a t t h e e f f e c t of t h e p r o p e l l e r , t h e n a c e l l e , and t h e i r DUCT CHOKE PLATE I

Figure 10. Isolated Propeller Apparatus - High Speed Tests

F i g u r e 11. P h o t o g r a p h o f Isolated P r o p e l l e r Model in Low Speed Wind T u n n e l / i S P A N h I S E S T A T I O N , 7 ) - 5 P.1 "i

/

I

F i g u r e 12. Model Wing P r e s s u r e I n s t r u m e n t a t i o n i n t e r a c t i o n on t h e aerodynamic loading of t h e wing could be determined.

Two chordwise rows were l o c a t e d on e i t h e r s i d e of t h e n a c e l l e , a t non- dimensional semispan s t a t i o n s 7 ) = 0.290, 0.328, 0.472, and 0.511. These p o s i t i o n s correspond t o d i s t a n c e s from t h e propfan c e n t e r l i n e of 85-percent and 55-percent of t h e p r o p e l l e r t i p r a d i u s on e i t h e r s i d e of t h e n a c e l l e . A f i f t h chordwise row was l o c a t e d n e a r t h e wing t i p a t 7 ) = 0.782 t o o b t a i n wing s u r f a c e p r e s s u r e d a t a o u t s i d e t h e i n f l u e n c e of t h e p r o p e l l e r and n a c e l l e . Each row c o n t a i n e d 18 o r i f i c e s on t h e upper s u r f a c e and 14 on t h e lower s u r f a c e . The chordwise l o c a t i o n s of t h e o r i f i c e s were c l o s e l y spaced a t t h e l e a d i n g edge t o c a p t u r e t h e h i g h pres- s u r e g r a d i e n t s and a l s o i n t h e 50-percent chordwise r e g i o n t o determine t h e shock l o c a t i o n a t t r a n s o n i c speeds.

Four a d d i t i o n a l spanwise o r i f i c e s t a t i o n s were l o c a t e d a t 1) = 0.231, 0.256, 0.545, and 0.569 t o h e l p d e f i n e wing loads. These were l i m i t e d t o e i g h t o r i f i c e s i n each row, two each on t h e upper and lower s u r f a c e s a t t h e l e a d i n g edge, and two each on t h e upper and lower s u r f a c e s a t t h e t r a i l i n g edge.

There were 24 s t a t i c p r e s s u r e o r i f i c e s on t h e n a c e l l e . E i g h t were l o c a t e d along t h e c e n t e r l i n e of t h e cowl on t o p of t h e n a c e l l e . On each s i d e of t h e n a c e l l e , e i g h t o r i f i c e s were l o c a t e d as shown i n F i g u r e 13.

F i g u r e 12 a l s o shows p r e s s u r e r a k e s j u s t a f t of t h e prop plane.

These rakes were used t o o b t a i n flow f i e l d d a t a on t h e semi-span model i n t h e NASA-Lewis 8-Ft x 6-Ft Supersonic Wind Tunnel as d i s c u s s e d i n S e c t i o n 3.4.6.

3.4.2 T o t a l P r e s s u r e Rakes I n l e t t o t a l p r e s s u r e r a k e s were l o c a t e d i n b o t h t h e PTA and Spey engine n a c e l l e s t o e v a l u a t e i n t e r n a l drag. The Spey i n l e t r a k e w a s a l s o designed t o measure flow d i s t o r t i o n , i f any, caused by t h e PTA i n s t a l l a - t i o n . On t h e PTA n a c e l l e , t h e flow-through duct on t o p of t h e n a c e l l e contained 5 t o t a l p r e s s u r e probes and 4 s t a t i c p r e s s u r e o r i f i c e s . The Spey n a c e l l e i n l e t had 1 2 t o t a l p r e s s u r e probes and 4 s t a t i c p r e s s u r e o r i f i c e s .

An e x t e r n a l rake c o n s i s t i n g of 29 t o t a l p r e s s u r e probes w a s mounted a t t h e e x i t p l a n e of t h e PTA n a c e l l e d u c t s o t h a t t h e j e t t h r u s t from t h e a i r t u r b i n e could be measured. The probes i n t h i s rake were arranged i n t h r e e h o r i z o n t a l rows and one v e r t i c a l row on t h e n a c e l l e e x i t c e n t e r l i n e .

3.4.3 Propfan Hub Balances The propfan hub balance ( F i g u r e was a non-rotating balance t h a t 14) measured f i v e components: axial f o r c e , normal f o r c e , p i t c h i n g moment, and t o a l i m i t e d accuracy, s i d e f o r c e and yawing moment. The balance was l o c a t e d between t h e d r i v e air-motor and t h e p r o p e l l e r hub. The balance c o n s i s t e d of two h a l f - c y l i n d e r s (one metric and one non-metric) connected by f o u r beams, and a s h a f t l o c a t e d by two b e a r i n g s i n t h e metric h a l f - c y l i n d e r of t h e balance. A f l e x u r e element t h a t t r a n s m i t s o n l y t o r q u e connected t h e s h a f t t o t h e d r i v e motor.

A f t e r an equipment f a i l u r e i n t h e high-speed t e s t s t h a t r e s u l t e d i n b a l a n c e f a i l u r e and l o s s of model hardware, a f a i l - s a f e r e d e s i g n was developed. To e n s u r e r e t e n t i o n i n t h e hub i n t h e e v e n t t h a t t h e beams o r 1.465 (57.667) 1.380 (54.333) FUSELAGE STATIONS 1.295 METER (INC11ES) (51.000) 1.211 (47.667) 1.126 (44.333) (40.444) 0.971

0.927 136.809) 0 . 935(36’[221 1

DENOTES S T A T I C (36.500)

PRESSURE PORT I

3 4

-

- - 5

I l 2

0.00338 22 23 Figure 13.

PTA Nacelle Pressure Instrumentation FLEXIBLE COUPLING RETENTION LUGS RETENTION RING T O METRIC SIDE Figure 14. Propfan Hub Balance t h e f l e x coupling were broken, a r e t e n t i o n r i n g w a s machined o n t o t h e m e t r F c h a l f of t h e b a l a n c e , and t h e two c y l i n d e r h a l v e s were made t o i n t e r l o c k by forming a 'T' shaped tongue on t h e non-metric p o r t i o n of t h e c y l i n d e r half and a corresponding c u t o u t on t h e metric c y l i n d e r h a l f w i t h a 0.102 c m (0.040 i n c h ) c l e a r a n c e allowed between t h e two. The hardware shown i n Figure 14 i n c l u d e s t h e elements of t h e redesign.

3.4.4 Low-Speed, Six-Component Force Balance Six-component f o r c e and moment d a t a were recorded on a l l r u n s u s i n g a 5.08 e m (2.0-inch) diameter i n t e r n a l balance. The balance was a NASA- Langley balance d e s i g n a t e d as Balance 748. It was a n orthogonal type such t h a t each l o a d component was measured by a s e p a r a t e s t r a i n gage bridge.

The f o r c e and moment l i m i t s of t h e balance were: N F f 8,007 N (1,800 l b ) f 2,224 N (500 l b ) AF P M f 791 m-N (7,000 in.-lb) f 4,448 N (1,000 l b ) SF yM f 339 m-N (3,000 in.-lb) RM f 452 m-N (4,000 in.-lb) The balance was a t t a c h e d t o a NASA-Langley o f f s e t a d a p t e r which mounts t o a s t i n g passing through t h e lower a f t fuselage. NASA-Langley provided t h e balance c a l i b r a t i o n and t h e c o r r e c t i o n s n e c e s s a r y t o account f o r t h e mechanical b r i d g i n g of t h e balance by t h e air supply l i n e and t h e i n s t r u m e n t a t i o n w i r e / t u b e bundle.

3.4.5 High-speed, Six-Component Force Balance The six-component balance used i n t h e high-speed tests was a 4.0-inch diameter i n t e r n a l balance b u i l t by Modern Machine and Tool Company and d e s i g n a t e d as Balance 984. This balance was a l s o an orthogonal type. The f o r c e and moment l i m i t s of t h e balance were: f 24,465 N (5,500 l b ) N F AF f 2,224 N (500 l b ) P M f 4,519 m-N (40,000 in.-lb) SF f 6,672 N (1,500 l b ) YM f 2,824 m-N (25,000 in.-lb) RM f 3,954 m-N (35,000 in.-lb) 3.4.6 Flow Survey Rakes To determine v e l o c i t i e s and flow a n g l e s i n and around t h e p r o p e l l e r p l a n e a s p e c i a l f l o w s u r v e y rake was employed. This r a k e , as shown i n Figure 15, contained f i v e , 5-hole probes one i n c h a p a r t on each of t w o arms. For surveys around t h e PTA n a c e l l e , t h e r a k e w a s mounted on a c o l l a r t h a t a t t a c h e d t o t h e nacelle j u s t behind t h e i n l e t cowl. T h i s c o l l a r allowed t h e r a k e t o be p o s i t i o n e d a t f o u r azimuthal l o c a t i o n s , and a t each l o c a t i o n t h e rake could be p o s i t i o n e d r a d i a l l y a t two l o c a t i o n s so t h a t p o i n t s one-half i n c h a p a r t could be obtained. The r a k e mount a l s o allowed a c e r t a i n amount of fore-and-aft p o s i t i o n i n g .

Design of t h e 5-hole probes was based on e x t e n s i v e e x p e r i e n c e w i t h such f l o w f i e l d measurements. Reference 1 p r e s e n t s t y p i c a l r e s u l t s from such instruments. E f f e c t i v e use of t h e s e probes, however, depends on a c a l i b r a t i o n t o d e f i n e t h e a n g u l a r s e n s i t i v i t y of t h e probes. Such a cali- b r a t i o n of t h e s e r a k e s w a s performed i n a small t r a n s o n i c wind t u n n e l a t Mach numbers t o 0.95 and i s d e s c r i b e d later, 3. 5 CALIBRATIONS 3.5.1 Propfan Airmotor C a l i b r a t i o n Because propfan t o r q u e could not be measured w i t h t h e hub balance, a c a l i b r a t i o n w a s performed t o determine a i r motor torque as f u n c t i o n s of d r i v e p r e s s u r e i n t o t h e motor and motor r o t a t i o n a l speed. The c a l i b r a t i o n was performed on a h y d r a u l i c dynamometer test r i g f o r two n a c e l l e config- u r a t i o n s : t h e PTA n a c e l l e and t h e i s o l a t e d n a c e l l e .

The c a l i b r a t i o n c u r v e s were of e x c e l l e n t q u a l i t y . It i s believed t h a t horsepower v a l u e s f o r t h e wind t u n n e l d a t a are a c c u r a t e w i t h i n about 0.447 kw (0.6 HP) or, a t maximum power, about 0.4-percent.

3.5.2 Flow Survey Probe C a l i b r a t i o n Details of a 5-hole probe are shown i n F i g u r e 16. Each probe con- s i s t e d of f o u r p i t o t t u b e s s o l d e r e d around a c e n t r a l p i t o t t u b e and a cone s t a t i c probe o f f s e t t o one s i d e . The c e n t r a l tube i n t h e bundle w a s f l a t - f a c e d , and t h e s i d e t u b e s were chamfered a t 45-degree angles. The r a k e s were c a l i b r a t e d i n t h e Lockheed-Georgia Compressible Flow Wind Tunnel (CFWT) by p l a c i n g each r a k e , i n t u r n , i n known flow c o n d i t i o n s and measuring t h e p r e s s u r e a t e a c h o r i f i c e on t h e rake. The f l o w c o n d i t i o n s were v a r i e d by changing t h e r o l l and p i t c h a n g l e s e t t i n g s of t h e r a k e and varying t h e t e s t s e c t i o n Mach number.

T h i s c a l i b r a t i o n and t h e 5-hole probe d a t a r e d u c t i o n procedures are d e s c r i b e d i n Appendix B.

3.6 W I N D TUNNEL AND MODEL INSTALLATIONS

3.6.1 High-speed Tests - Langley 16-Ft T r a n s o n i c Aerodynamics Wind Tunnel

The NASA-Langley 16-Ft Transonic Aerodynamics Wind Tunnel i s a s i n g l e r e t u r n atmospheric wind t u n n e l w i t h Mach number range from 0.2 t o 1.3.

The s l o t t e d octagonal t e s t s e c t i o n nominally measures 4.72m (15.5 f e e t ) a c r o s s t h e f l a t s and has a u s a b l e test s e c t i o n l e n g t h of 6.71m ( 2 2 feet).

Figure 15. Flow S u r v e y Rake g-, I I I I !- I CONE S T A T I C F I V E . IlOLE PROBE Figure 16. 5-Hole Probe

*-

The PTA model w a s sting-mounted i n t h e t u n n e l as shown i n Figure I.

Model blockage based on maximum f r o n t a l area was less t h a n 1 percent.

Reynolds numbers, based on MAC ranged from about 1.8 m i l l i o n a t Mach 0.2 t o about 5 m i l l i o n a t Mach 0.8.

Temperature i n t h e 16-ft Transonic Aerodynamics Wind Tunnel normally i n c r e a s e d w i t h o p e r a t i n g t i m e s o t h a t t h e model seldom, i f e v e r , a r r i v e d a t thermal equilibrium. Unfortunately, t h e balance and a i r bypass system was s e n s i t i v e i n t h e d r a g d i r e c t i o n t o thermal g r a d i e n t s . Even though an a t t e m p t w a s made t o measure model, balance, and bypass l i n e temperatures and compensate f o r t h e temperature g r a d i e n t s , t h i s w a s n o t completely s u c c e s s f u l . E r r o r s i n axial f o r c e measurements f o r t h e powered runs were, t h e r e f o r e , h i g h e r t h a n d e s i r e d . With t h i s e x c e p t i o n , t h e flow q u a l i t y i n t h i s f a c i l i t y was good, and d a t a from t h e tests were of h i g h q u a l i t y .

3.6.2 Low-Speed T e s t s - Langley 4M x 7M Subsonic Wind Tunnel

The NASA-Langley 4 M x 7M Subsonic Wind Tunnel i s a s i n g l e r e t u r n subsonic wind t u n n e l w i t h Mach numbers t o about 0.3. The t e s t s e c t i o n i s 4 m (14.5-feet) high and 7 m (21.75-feet) wide. The P T A model was s t i n g - mounted i n t h e test s e c t i o n a s shown i n F i g u r e 4. Tests were run p r i m a r i l y a t M = 0.165 and q = 1915 N / m 2 (40 p s f ) although a few runs were made a t M = 0.2 and q = 2873 N / m 2 (60 p s f ) . Reynolds numbers based on mean aerodynamic chord ranged from 1.6 x lo6 t o 2 x 10 .

Although t h e b a l a n c e / a i r bypass system w a s d i f f e r e n t from t h a t used i n t h e high-speed tests, problems were a g a i n encountered w i t h measuring power-on drag. Power-off d r a g was r e l i a b l e as were a l l o t h e r f o r c e s and measurements w i t h power on, but t h e power-on d r a g d a t a were not good-- a p p a r e n t l y due t o i n t e r f e r e n c e between t h e s t i n g and t h e a i r bypass l i n e .

3.6.3 High-speed Flow Survey Tests - L e w i s 8-Ft x 6-Ft Supersonic Wind

Tunnel The NASA-Lewis 8-Ft x 6-Ft Supersonic Wind Tunnel, i n i t s aerodynamic c y c l e , is operated as a c l o s e d system w i t h d r y a i r added as r e q u i r e d t o maintain t h e d e s i r e d t u n n e l c o n d i t i o n . The t u n n e l is capable of o p e r a t i n g i n t h e Mach number range from 0.36 t o 2.0, b u t f o r t h e s e tests was o p e r a t e d a t Mach numbers from 0.6 t o 0.85. Reynolds number, based on MAC, ranged from about 4.7 x lo6 t o 5.6 x lo6.

The test s e c t i o n is 2.44m ( 8 f e e t ) h i g h , 1.83m ( 6 f e e t ) wide, and 7.16m (23.5 f e e t ) long. It is p e r f o r a t e d on a l l f o u r w a l l s t o provide approximately 6-percent p o r o s i t y .

The h a l f - f u s e l a g e , semispan PTA model was mounted on a f l a t p l a t e which i n t u r n w a s i n s t a l l e d along one s i d e w a l l as shown i n F i g u r e 6. The s i d e w a l l p l a t e i s d i s p l a c e d 15.24 c m (6 i n c h e s ) from t h e t u n n e l wall t o minimize w a l l boundary l a y e r contamination of flow around t h e semispan model.

Blockage of t h e model i n t h e tunnel was 1.87 p e r c e n t of t h e t u n n e l c r o s s - s e c t i o n a l area.

4 . 0 TEST PROCEDURES I n t h e wind t u n n e l environment, tests are g e n e r a l l y conducted w i t h Reynolds numbers about one o r d e r of magnitude lower t h a n i n f u l l scale f l i g h t . The major impact of t h i s Reynolds number d i s p a r i t y is on v i s c o u s phenomena l i k e boundary l a y e r growth, t r a n s i t i o n from laminar t o t u r b u l e n t flow, and boundary l a y e r s e p a r a t i o n . I n f u l l - s c a l e f l i g h t , the boundary l a y e r i s t u r b u l e n t over most of t h e a i r c r a f t s u r f a c e ; on t h e models, how- e v e r , t h e boundary l a y e r w i l l u s u a l l y be l a m i n a r u n l e s s it is a r t i f i c i a l l y t r i p p e d . S p e c i a l care must b e e x e r c i s e d i n t h e wind t u n n e l t o a t t a i n t h e boundary l a y e r c o n d i t i o n s that g i v e proper s i m u l a t i o n of f l i g h t .

For high-speed wind t u n n e l tests, e x p e r i e n c e h a s shown t h a t i t i s important t o produce t h e t h i n n e s t p o s s i b l e t u r b u l e n t boundary l a y e r on wing s u r f a c e s . I f t h e boundary l a y e r is t h i c k e r than f o r f u l l - s c a l e simu- l a t i o n , s t r o n g a d v e r s e p r e s s u r e g r a d i e n t s such as t h o s e a s s o c i a t e d with t r a n s o n i c shock waves may cause t h e boundary l a y e r t o separate ahead of t h e p o s i t i o n where it would s e p a r a t e i n f l i g h t . I n some cases, t h i s h a s caused wind t u n n e l f o r c e s and moments t o be much d i f f e r e n t from t h o s e experienced i n f l i g h t . It i s something of an a r t , t h e r e f o r e , t o t r i p t h e l a m i n a r boundary l a y e r a t t h e p l a c e where t h e b e s t s i m u l a t i o n of f l i g h t c h a r a c t e r i s t i c s i s achieved.

I n t h e wind t u n n e l , i f t h e boundary l a y e r i s t r i p p e d t o o f a r forward, it w i l l t h i c k e n d i s p r o p o r t i o n a t e l y i n t h e low Reynolds number environment and be t o o t h i c k i n t h e a f t r e g i o n s of t h e wing. Standard procedure, t h e r e f o r e , is t o p l a c e t h e boundary l a y e r t r i p as f a r a f t as p o s s i b l e and s t i l l be a s s u r e d t h a t t h e t u r b u l e n t boundary l a y e r will be well-developed i n t h e r e g i o n s where it w i l l be s u b j e c t e d t o s t r o n g a d v e r s e p r e s s u r e g r a d i e n t s . Usually, t h e test personnel a t a g i v e n f a c i l i t y w i l l have had enough e x p e r i e n c e w i t h models of a g i v e n type t o have a good " f e e l " f o r l o c a t i o n of t r a n s i t i o n t r i p s , and c o n s i d e r a b l e reliance i s placed on t h e i r judgement .

I n low-speed wind t u n n e l tests, t h e r e are no s t r o n g shock-wave- induced adverse p r e s s u r e g r a d i e n t s t o contend w i t h , and g e n e r a l l y i n low-speed wind t u n n e l s , t h e d a t a are much less s e n s i t i v e t o boundary l a y e r s i m u l a t i o n procedures. Again, however, t h e personnel of a g i v e n f a c i l i t y are u s u a l l y t h e b e s t s o u r c e of guidance i n t h i s matter.

These and o t h e r f a c t o r s p e r t i n e n t t o t h e p r o c e s s of o b t a i n i n g good f l i g h t d a t a i n PTA scale model wind t u n n e l tests are d i s c u s s e d i n t h e f o l l o w i n g s e c t i o n s .

4.1 HIGH-SPEED TESTS - 16-FT TRANSONIC AERODYNAMICS WIND TUNNEL

High-speed tests were performed w i t h th,e G I 1 model and w i t h the model c o n f i g u r e d i n t h e PTA c o n f i g u r a t i o n . I n e a c h case, runs were made w i t h t a i l on and o f f . C o n f i g u r a t i o n v a r i a b l e s on t h e PTA model i n c l u d e d t h e e f f e c t s of s p o i l e r , a i l e r o n , and rudder d e f l e c t i o n . Data were a l s o o b t a i n e d f o r f e a t h e r e d p r o p e l l e r , windmilling p r o p e l l e r , and powered pro- p e l l e r . P r o p e l l e r motor power l i m i t a t i o n s were such, however, t h a t i n t h e s e high-speed tests, t h r u s t c o e f f i c i e n t s were l i m i t e d t o about 0.10.

A series of tests were run t o d e t e r m i n e t h e e f f e c t s of boundary l a y e r t r i p p i n g technique. F i g u r e s 17 and 18 show r e s u l t s a t Mach 0.85 of P Q - DEGREES Figure 17. Effects of Transition Fix on Lift Figure 18. Effects of Transition Fix on Drag s e v e r a l t r i p l o c a t i o n s on l i f t curves and d r a g polars. T r a n s i t i o n was f i x e d by g l u i n g small g l a s s s p h e r e s ( B a l l o t i n i b a l l s ) i n a narrow s t r i p on t h e wing s u r f a c e . Fixing t r a n s i t i o n at 10-percent chord r e s u l t e d i n h i g h e r d r a g than a t t h e 30-percent chord l o c a t i o n and a l s o r e s u l t e d i n a break o f f of t h e l i f t curve s l o p e a t a lower angle of a t t a c k . T h i s i m p l i e s t h a t t h e boundary l a y e r was u n n e c e s s a r i l y thickened by t r i p p i n g a t t h e 10-percent chord l o c a t i o n . Furthermore, s i n c e t h e r e is no i n d i c a t i o n t h a t t r i p p i n g a t 30-percent chord produced any premature s e p a r a t i o n or unnecessary d r a g , t h e s e d a t a i n d i c a t e that t h e 30-percent l o c a t i o n is p r e f e r a b l e . Most of t h e high-speed wind t u n n e l d a t a , t h e r e f o r e , were o b t a i n e d with t r a n s i t i o n f i x e d a t t h e 30-percent chord l o c a t i o n .

T e s t c o n d i t i o n s included Mach numbers of 0.4, 0 . 7 , 0.8, and 0.85, and a n g l e s of a t t a c k ranging from -2 degrees up t o t h e angle of a t t a c k f o r b u f f e t onset. Two p r o p e l l e r advance r a t i o s were set f o r each combination of Mach number and blade p i t c h angle. Table I summarizes t h e t e s t con- d i t i o n s .

4.2 LOW-SPEED TESTS - 4M X 7M SUBSONIC W I N D TUNNEL

Low-speed t e s t s were performed i n a manner s i m i l a r t o t h a t of t h e high-speed tests. The b a s i c GI1 model was t e s t e d w i t h t a i l off and on; and w i t h runs t o measure t h e e f f e c t s of f l a p s , s p o i l e r s , a i l e r o n , rudder, and e l e v a t o r . For t h e PTA c o n f i g u r a t i o n , i n a d d i t i o n t o t h e above, v a r i a b l e s included t i p booms, t h e a c o u s t i c boom, PTA n a c e l l e i n c i d e n c e , and propfan w i t h and without power. The propfan was a l s o t e s t e d on t h e i s o l a t e d n a c e l l e . Flow survey tests were made on t h e model i n t h e p r o p e l l e r plane (prop-off) and on t h e i s o l a t e d n a c e l l e r i g both i n t h e p r o p e l l e r plane and behind t h e powered p r o p e l l e r .

A l l of t h e model tests i n t h e low-speed t u n n e l were made w i t h t r a n s i - t i o n f r e e on t h e advice of t h e f a c i l i t y t e s t personnel. Most of t h e t e s t s were r u n a t a Mach number of 0.165. Model a n g l e s of a t t a c k ranged from -2 degrees t o a t l e a s t t h e s t a l l c o n d i t i o n , and, when model v i b r a t i o n was not e x c e s s i v e , t o s e v e r a l d e g r e e s beyond stall. Some tests were performed a t Mach 0.2 w i t h p r o p e l l e r powered i n o r d e r t o g e t a more a p p r o p r i a t e range of advance r a t i o s . The same was t r u e f o r t h e i s o l a t e d p r o p e l l e r tests.

Two b l a d e angle s e t t i n g s were used f o r a l l of t h e powered p r o p e l l e r tests.

Run c o n d i t i o n s f o r t h e low-speed tests are summarized i n Table 11.

4 . 3 HIGH-SPEED FLOW SURVEY TESTS - 8-FT X 6-FT SUPERSONIC WIRD TUNNEL

For t h e high-speed flow survey t e s t s on t h e semispan model, a l l of t h e flow surveys were made i n t h e p r o p e l l e r p l a n e w i t h t h e p r o p e l l e r off.

Mach numbers were v a r i e d from 0.60 t o 0.85; n a c e l l e tilt a n g l e s t e s t e d were -3, -1, and +2 d e g r e e s ; and model a n g l e of a t t a c k ranged from -2 t o +6 degrees. I n a d d i t i o n t o flow survey rake p r e s s u r e s , p r e s s u r e d i s t r i b u - t i o n s on t h e wing were measured t o compare w i t h those from o t h e r scale tests.

Run c o n d i t i o n s for t h e s e tests a r e summarized i n Table 111.

TABLE 1 .

HIGH SPEED TEST CONDITIONS (a) Bo Power On Propfan MACH NO.

- - - - -

BL

.7a .a . a2 .as

I CONFIGURATION PROP TRANSITION . 4 .7

= - = =

-

--

X X

G I 1 - U p r i g h t B X X

--

B X X X X

G I 1 - I n v e r t e d

--

B X X X

G I 1 - T a i l On X

X X X F r e e X A X X X X X X

--

X X

G I 1 - T a i l O f f B X X

X

PTA - T a i l O f f O f f B X X X

X

PTA - T a i l On O f f B X X X

X X

PTA - W/O A c o u s t i c Boom O f f B X X

X X X X

PTA - No Booms O f f B X X

X X X X X F r e e X A X X X X X X X X X X PTA t o LEX O f f A X X X X X X PTA + LEX + Nacelle Bump O f f A

- - - -

- -

T r a n s i t i o n

A - F i x e d a t 10% Chord

B - F i x e d a t 30% Chord

TABLE I .

HIGH SPEED TEST CONDITIONS (b) Powered Propfan

- - - -

PROPFAN BL MACH ADVANCE 6 , 6 A NO.

RATIO, J TRANS IT I O N CONFIGURATION

- - -

-

~ ~ ~~ .4 B

PTA - Tail On 1.7

. 7 2 . 9 B .7 2 . 9 B . 8 3 . 3 B B . 8 3.06 Free . 8 3.06 .4 +5 ' 1.8 B Rudder Deflection . 7 +5 a 2.913.2 B . 8 +5 O 3.21 3.37 B B . 4 -5 * 1 . 8 . 7 -5 * 2 . 9 / 3 . 2 B . 8 - 5 " 3 . 2 / 3 . 3 7 B B . 4 -10" Aileron Study 1.71 1.8 B . 7 -10' 2 . 9 1 3 . 2 -10' B .8 3.2/3.37 . 4 - 1 0 ' B Spoiler Study 1.711.8 .7 -10' B 2.913.2 -10" B . 8 3 . 6 - 1 0 ' B .8 3.2/3.37 .4 -10' -10' 1.8 B Aileron h Spoiler .7 -10" - 1 0 ' B 2,913.2 -10' - 1 0 ' B .a 3.213.37 .4 1.8 B PTA & LEX .7 2.913.0613.2 B . 8 3.313.4 B .8 3.213.37 B

-

TRANSITION

A - Fixed a t 10% Chord

B - Fixed a t 30% Chord

3 f '0.4

c - -

x x x X X x x X x x x x x X x x

-

- N

N a a n a a a m a l m m W W B

v) ; i ; w L

.

n El a + a

. -

Q Y n In m C N II a I m

Y -

A= Y Y C t k C b b t P a I

- - a 8 Y

- -

a -

I - W h 8 TABLE 111.

HIGH SPEED FLOW SURVEY TEST CONDITIONS .

a NT COLLAR

__+_

-1 " -2" t o +6"

- 6 - - 8 5 On

+2 " -2" to +6" On

- 6 - e 8 5

+2 O -2" to +6" Off 4 Horizontal Rake Positions and 4 Vertical Rake Positions for Each Point 4 . 4 FORCE MEASUREMENT TARES, CALIBRATIONS, AND CORRECTIONS I n p r e p a r a t i o n f o r t h e PTA wind t u n n e l tests, s e v e r a l i m p o r t a n t cali- b r a t i o n s were performed. The f i r s t of t h e s e was t h e p r o p e l l e r d r i v e motor c a l i b r a t i o n d e s c r i b e d earlier. T h i s c a l i b r a t i o n was needed because t h e r e were no p r o v i s i o n s on t h e model f o r measuring p r o p e l l e r torque.

The exhaust from t h e p r o p e l l e r d r i v e motor provided a s i g n i f i c a n t t h r u s t component t h a t r e q u i r e d s p e c i a l c a l i b r a t i o n s i n t h e wind tunnel.

F i r s t , a s t a t i c tare c a l i b r a t i o n was performed w i t h t h e PTA n a c e l l e r e p l a c e d by a c a l i b r a t e d j e t nozzle. T h i s test determined t h e i n t e r - a c t i o n s between nozzle t h r u s t and o t h e r f o r c e s and moments measured d u r i n g wind t u n n e l tests. A second s t a t i c tare w a s performed w i t h t h e PTA n a c e l l e on t h e wing and t h e d r i v e motor pinned t o p r e v e n t r o t a t i o n i n o r d e r t o determine t h e r e s i d u a l j e t t h r u s t of t h e PTA n a c e l l e . F i n a l l y , w i t h wind on, t h e same model set-up w a s t e s t e d w i t h and w i t h o u t j e t flow t o provide a j e t t h r u s t c a l i b r a t i o n t h a t included i n t e r f e r e n c e e f f e c t s w i t h t h e t u n n e l flow.

I n t h e powered wind t u n n e l tests, p r o p e l l e r t h r u s t was measured w i t h t h e p r o p e l l e r hub b a l a n c e between t h e d r i v e motor and t h e p r o p e l l e r . To o b t a i n d r a g , t h i s t h r u s t w a s c o r r e c t e d f o r base and forward s u r f a c e pres- s u r e f o r c e s on t h e hub and added t o f o r c e b a l a n c e measurements.

A s mentioned e a r l i e r , t h e f o r c e b a l a n c e used i n t h e high-speed wind t u n n e l w a s s e n s i t i v e t o temperature e f f e c t s i n t h e d r a g d i r e c t i o n . This s e n s i t i v i t y o r i g i n a t e d i n t h e h i g h p r e s s u r e a i r l i n e s ( f o r p r o p e l l e r d r i v e ) t h a t passed around t h e balance. The bellows used i n t h e s e l i n e s were n o t s u f f i c i e n t l y f l e x i b l e t o compensate f o r thermal expansion f o r c e s a t h i g h p r e s s u r e s . The problem was aggravated by t h e o p e r a t i o n a l charac- t e r i s t i c of t h e wind t u n n e l t h a t r e s u l t e d i n a c o n t i n u o u s l y i n c r e a s i n g a i r temperature w i t h i n c r e a s i n g run t i m e . An a t t e m p t was made t o c a l i b r a t e t h i s f o r c e by i n s t r u m e n t i n g t h e balance w i t h a number of thermocouples.

T h i s e f f o r t was u n s u c c e s s f u l , however, s o t h e powered d r a g d a t a f o r t h e high-speed tests was n o t r e l i a b l e . This t h e r m a l expansion e f f e c t d i d n o t , however, degrade o t h e r f o r c e measurements ( l i f t , s i d e f o r c e ) s i n c e t h e bypass l i n e s were s u f f i c i e n t l y f l e x i b l e i n a l l d i r e c t i o n s e x c e p t a l o n g t h e b a l a n c e a x i s .

I n t h e low-speed wind t u n n e l , a d i f f e r e n t balance system was used but a g a i n t h e d r a g measurements w i t h t h e propfan powered were e r r a t i c and g e n e r a l l y u n r e l i a b l e . These d i s a p p o i n t i n g r e s u l t s h i g h l i g h t t h e d i f f i - c u l t i e s involved i n o b t a i n i n g good d r a g d a t a when h i g h p r e s s u r e a i r l i n e s must bypass a f o r c e b a l a n c e i n a r e l a t i v e l y small model. T i m e c o n s t r a i n t s i n t h i s t e s t program d i d n o t permit t h e development needed t o s o l v e t h e s e measurement problems.

3 ’ 5.0 DATA ACCURACY The following examples provide estimates of d a t a accuracy i n t h e measurement and c a l i b r a t i o n of test parameters and performance coef- f i c i e n t s . A l l p r e s s u r e s are normalized with r e f e r e n c e t o s t a n d a r d atmospheric p r e s s u r e , 101,325 N/mz (2,116 psf).

Frees tream Mach Number Y- 1

-

- I]>’

Assume Mo = 0.6094, y = 1.4 S t a t i c P r e s s u r e Error = T o t a l P r e s s u r e Error = k.002362 L e t Ho = 1.2141 and Po = 0.9448 o Max/Min Error Analysis With t h e above p r e s s u r e s : = 0.6149 MO max = 0.6038 MO m i n A M = f0.9% o RMS Error Analysis e r r o r aMO aMO - =

1.334 1.037; - =

aHO aPO

RMS E r r o r = [(1.037*.002362)2 + (1.334*.002362)2]

= .00399 o r 0.65% P Freestream Dynamic Pressure For p = 0,9448 and Mo * 0.6094 qo = 0.2456 o Max/Min Error Analysis For the pressure errors c i t e d e a r l i e r (f.002362) = .2507 QO max = .2405 min o UMS Error Analysis RMS Error = error

- a q O = 0.260; - aqO = 0.806

aPO aMO RMS E r r o r = [(0.26*.002362)2 + (0.806*.0054) 2 + ] = 0.004396 or 1.79% Model S t a t i c Pressure Coefficient

P - Po

0 c p = = 0.9448 f .002362 Assume : PO p = 1,1810 f .00142 = 0.2456 f .004724 QO o Max/Min Error Analysis C = 0.962 'nomina 1 C = 0.998 Pmax c = 0.927 Pmin AC = f 3.7% P o RMS Error Analysis

ac ac ac

4.072; 2 = -4.072; 2 = -3.915 a P aPO a q O

R M S Error = [(4.072*.00142)2 + (-4.072*.002362)

2 +

+ (-3.915*.004724) ]

= 0.0216 o r 2.24% The measurement s y s t e m f o r t h e flow survey r a k e s used t h e same t r a n s d u c e r , so t h a t p r e s s u r e c o e f f i c i e n t s were s u b j e c t t o t h e same e r r o r s .

Balance U n c e r t a i n t i e s The balance a c c u r a c i e s f o r each component are g i v e n i n Table TV.

Aerodynamic C o e f f i c i e n t U n c e r t a i n t i e s The u n c e r t a i n t i e s i n t h e balance when combined w i t h u n c e r t a i n t i e s i n t h e f r e e s t r e a m dynamic p r e s s u r e g i v e t h e following u n c e r t a i n t i e s i n l i f t c o e f f i c i e n t and d r a g c o e f f i c i e n t .

Assume a c r u i s e c o n d i t i o n of Mach = 0.8, f r e e s t r e a m dynamic p r e s s u r e f 1 2 p s f ) , and d r a g f o r c e of D = 1010 of q = 29,686 f 574.6 N / m 2 (620 f 1 1 N ( 2 2 7 f 2 . 5 l b ) . The maximum u n c e r t a i n t y i n d r a g c o e f f i c i e n t would be 0.0004 o r 4 counts (about 1 . 1 p e r c e n t of c r u i s e d r a g c o e f f i c i e n t ) . For a l i f t f o r c e , L = 9261 f 1 2 2 N (2,082 f 27.5 l b ) t h e maximum u n c e r t a i n t y i n l i f t c o e f f i c i e n t would be 0.004 (about 1 . 4 p e r c e n t of c r u i s e l i f t c o e f f i c i e n t ) .

TABU IV.

BALANCE ACCURACIES L O W SPEED BALANCE L O W SPEED BALANCE H I G H SPEED COMPONENT LIMITS UNCERTAINTIES LIMITS UNCERTAINTIES f122.0 N (27.5 l b ) f l l . O N (2.5 l b ) f25.0 m-N (200.0 i n - l b ) k33.0 N (7.5 l b ) f14.0 m-N Yawing Moment 339 m-N f 1 . 7 m-N 2,824 m - N (125.0 i n - l b ) (3,000 i n - l b ) (15.0 i n - l b ) (25,000 i n - l b ) R o l l i n g Moment 452 m-N f2.3 m-N 3,954 m-N f20.0 m-N (4,000 i n - l b ) (20.0 i n - l b ) (35,000 i n - l b ) (175.0 i n - l b ) 6 . 0 ANALYTICAL PREDICTIONS The CATIA code (Reference 2 ) was used to manipulate and store configuration geometry for the models.

This code allowed simple and rapid configuration changes to be input to the flow prediction codes. A typical panel model of a PTA configuration is shown in Figure 19.

Well developed analytical codes were available to predict aerodynamic and stability and control characteristics of the PTA model. The basic code was QUADPAN (Reference 3)--an advanced, low order, three-dimensional panel code that has been widely used in aircraft design studies. QUADPAN, while not effective in supersonic flow regions, does incorporate compressibility corrections so that it is reasonably accurate for transonic flows if imbedded supersonic regions are not dominant. A propeller performance code, PROPVRTX, was used to predict slipstream properties. This code and its validation are described in Reference 4 .

PROPVRTX was interfaced with QUADPAN by restating QUADPAN surface boundary conditions to include velocity perturbations calculated with PROPVRTX, and then correcting surface pressures washed by the slipstream for the pressure rise across the propeller disc. The efficacy of this technique was demonstrated in Reference 5.

The proper level of slipstream thrust was set by scaling axial and tangential velocity distributions to correspond to the desired propfan Cp and J values. The basic slipstream model was derived from flow survey tests reported in Reference 6 .

Special consideration was required and given to: treatment of the aircrafr wake, simulation of the PTA nacelle exhaust, and accounting for internal flow through the engine nacelle inlets. These and other aspects of the analytical predictions are discussed in Appendix C .

Y \ G Figure 19. QUADPAN Panel Model of PTA Configuration 7.0 RESULTS AND DISCUSSION 7.1 PERFORMANCE DATA The n a t u r e of t h e PTA a i r c r a f t and t h e PTA program o b j e c t i v e s re s u l t ed i n t h e following emphases i n t h e wind tunnel program: 0 N o powered wind t u n n e l tests w i t h f l a p s extended were performed because propfan o p e r a t i o n w i t h f l a p s extended was p r o h i b i t e d on t h e a i r c r a f t .

0 A heavy emphasis w a s placed on high-speed d r a g because t h e major a i r c r a f t performance concern w a s on attainment of Mach 0.8 d e s i g n c r u i s e speed and speeds t o Mach 0.85, i f possible.

0 A s t r o n g emphasis w a s placed on s t a b i l i t y and c o n t r o l a t low speeds because low a l t i t u d e f l y o v e r s for n o i s e measurements were required.

The PTA model was complicated by t h e powered n a c e l l e which had a s e p a r a t e f o r c e balance f o r p r o p e l l e r f o r c e s and moments, and an a i r exhaust nozzle t h a t produced s i g n i f i c a n t t h r u s t when t h e a i r motor was o p e r a t i n g t o d r i v e t h e propfan. A d e s c r i p t i o n of model f o r c e bookkeeping methods and c o n s i d e r a t i o n s f o r d r a g s c a l i n g a r e given i n Appendix D. Drag c o e f f i c i e n t s presented i n t h i s d i s c u s s i o n , u n l e s s o t h e r w i s e noted, a r e defined a s follows:

CD = Model Balance Drag + P r o p e l l e r Thrust i Powered Nacelle J e t Thrust

7.1.1 L i f t / P i t c h i n R Moment 7.1-1-1 Data C o r r e l a t i o n With Theory The b a s i c t o o l used f o r aerodynamic a n a l y s i s i n t h e PTA Program w a s QUADPAN. It was f i r s t v a l i d a t e d by comparing p r e d i c t e d performance w i t h G I 1 a i r c r a f t d a t a . Such a comparison is shown i n Figure 20. It published can be seen t h a t QUADPAN p r e d i c t s t h e C L v e r s u s a ! curve extremely w e l l , while a good p r e d i c t i o n i s obtained f o r t h e C L v e r s u s C , curve. F i g u r e 21 shows a comparison of r e s u l t s from t h e QUADPAN a n a l y t i c a l code w i t h wind t u n n e l d a t a f o r the G I 1 model. Again, t h e r e i s e x c e l l e n t agreement between theory and experiment f o r t h e CL - C Y curve and good agreement f o r CL-C, curve. The divergence of d a t a and p r e d i c t e d c u r v e s f o r p i t c h i n g t h e moment i m p l i e s t h a t t h e n e u t r a l p o i n t f o r t h e QUADPAN p r e d i c t i o n w a s s l i g h t l y a f t of t h a t measured i n t h e wind tunnel.

Even The c o r r e l a t i o n shown i n Figure 2 1 i s f o r a Mach number of 0 . 4 .

though QUADPAN is b a s i c a l l y an incompressible code, i t is s t i l l q u i t e e f f e c t i v e a t h i g h e r Mach numbers a s shown i n Figure 22. I n t h i s f i g u r e , d a t a and p r e d i c t i o n s a r e shown f o r t h e PTA model a t Mach 0.7. It can be seen t h a t t h e l i f t curve i s not p r e d i c t e d q u i t e so p r e c i s e l y as f o r t h e s i m p l e r c o n f i g u r a t i o n a t . t h e lower Mach number, but t h e disagreement is

e -

M / I I I I - 0 . 0 4 - 0 . 0 8 - 0 . 1 2 -0.lG Q c -.

Figure 20. Comparison o f QUADPAN Predictions with GI1 Data GI1 FORCE D A T A A T MACH 0.4 0 PREDICTED QUADPAN A M E A S U R E D LRC 16' TUNNEL 0.04 0.00 -0.04 -0.08 PITCHING MOMENT, Cm ANGLE OF A T T A C K , U Lift/Pitching Moment Correlation with Theory - GI1 Figure 2 1 .

PTA FORCE D A T A AT MACH 0.7 0 PREDICTED QUADPAN A MEASURED LRC 16' TUNNEL ANGLE OF A T T A C K , U PITCHING MOMENT, Cm Figure 2 2 . Lift/Pitching Moment Correlation w i t h Theory - PTA b e l i e v e d t o be due more t o t h e a d d i t i o n a l complexity of t h e PTA model t h a n t o t h e h i g h e r Mach number. Again, t h e p i t c h i n g moment curve is n o t p r e - d i c t e d as w e l l a s t h e l i f t c u r v e , b u t t h e agreement is s t i l l good.

The e x c e l l e n t c o r r e l a t i o n s of F i g u r e s 20 through 22 g i v e confidence i n t h e v a l i d i t y of t h e wind t u n n e l d a t a and t h e a b i l i t y of t h e QUADPAN code a s a t o o l f o r i n t e r p r e t a t i o n and e x t r a p o l a t i o n of data.

7.1.1.2 C o n f i g u r a t i o n Buildup Figure 23 shows t h e manner i n which buildup of t h e model from t h e GI1 t o t h e PTA c o n f i g u r a t i o n a f f e c t s l i f t . The d a t a are shown f o r Mach 0.8 where t h e e f f e c t s of t i p and a c o u s t i c booms, i f any, would be expected t o be e v i d e n t . The d a t a of F i g u r e 2 3 show t h a t o n l y t h e a d d i t i o n of t h e n a c e l l e t o t h e G I 1 wing h a s any n o t i c e a b l e e f f e c t on t h e l i f t c h a r a c t e r - istics. The PTA n a c e l l e i n c r e a s e s c o n f i g u r a t i o n l i f t s l i g h t l y a t low a n g l e s of a t t a c k , and d e c r e a s e s l i f t s l i g h t l y a t a n g l e s of a t t a c k above 3 degrees.

F i g u r e 24 shows t h a t t h e t r e n d observed a t Mach 0.8 appeared a t lower Mach numbers a l s o . A t Mach 0.4 and 0.17 t h e PTA n a c e l l e i n c r e a s e d l i f t a t low a n g l e s of a t t a c k b u t r e s u l t e d i n a s l i g h t l o s s of l i f t a t t h e higher angles.

The effects of f l a p s on l i f t are shown f o r t h e GI1 model i n Figure 25 and f o r t h e PTA model i n Figure 26. The adverse e f f e c t of t h e PTA config- u r a t i o n on maximum l i f t c o e f f i c i e n t i s e v i d e n t f o r takeoff f l a p s but i s n o t s e e n f o r landing f l a p s .

7.1.1.3 E f f e c t s of Mach Number The e f f e c t s of Mach number on t h e l i f t c h a r a c t e r i s t i c s of t h e PTA model a r e summarized i n Figure 27. I n c r e a s i n g Mach number i n c r e a s e d a !

f o r z e r o l i f t , i n c r e a s e d t h e s l o p e of t h e l i f t c u r v e , and decreased l i f t a t t h e h i g h e r a n g l e s of a t t a c k .

7.1.1.4 E f f e c t s of t h e Propfan T h r u s t c o e f f i c i e n t s a v a i l a b l e f o r t h e high-speed tests were low, and t h e p r o p f a n e f f e c t s on l i f t t h a t were measured were i n s i g n i f i c a n t . A t low s p e e d s , however, a wide range of t h r u s t c o e f f i c i e n t v a l u e s were a v a i l a b l e , and s i g n i f i c a n t e f f e c t s were observed. These w i l l be d i s c u s s e d f u l l y i n S e c t i o n 7 . 3 STABILITY AND CONTROL. S u f f i c e i t t o say i n t h i s s e c t i o n t h a t both l i f t and l i f t curve s l o p e i n c r e a s e w i t h i n c r e a s i n g propfan t h r u s t w i t h most of t h i s i n c r e a s e i n l i f t a t t r i b u t e d t o t h e e f f e c t of i n c r e a s e d v e l o c i t i e s i n t h e s l i p s t r e a m .

Drag p o l a r s f o r t h e G I 1 model i n t h e low-speed wind t u n n e l are shown i n F i g u r e 28. The s e v e r a l curves show t h e model i n t h e t a i l - o f f , v e r t i c a l - t a i l - o n l y , and t a i l - o n c o n f i g u r a t i o n s . Also p l o t t e d i n Figure 28 i s a d r a g p o l a r from published G I 1 a i r c r a f t f l i g h t performance d a t a a t Mach 0.20. I n t h e range of l i f t c o e f f i c i e n t s from 0.1 t o 0 . 6 , i t can be seen t h a t t h e low-speed wind t u n n e l d a t a g i v e G I 1 d r a g about 10 t o 40 MACH 0.80 CONFIGURATION 1 . 0 4 0 GI I GI1 + PTA NACELLE P T A LESS ACOUSTIC BOOM A 0 . 8

* PTA

a. 1 0 .

a ANGLE OF A T T A C K , a - 0 . 2 J Lift Curves for PTA Buildup Figure 2 3 .

Y

*-

Figure 24. PTA Lift Coefficients at Several Mach Numbers

*-

Figure 2 5 . Effect of Flaps on Lift - GI1

- -

Figure 2 6 . Effect of Flaps on Lift - PTA 1.0

u -I 0.8!

.

I- Z 0.6 //

W

-

MACH

u

-

LL 0 . 4 ~ u , 0 0.40 W 0.70 0 . 8 0

W A

0 . 8 5

*

ANGLE OF A T T A C K , U Figure 27. Effect of Mach Number on PTA Lift 1 . 2 1.0 0.8 0.6 cL 0.4 0.2 -0.2 -0.4 Figure 28. Drag Polars for GI1 counts h i g h e r than t h e f l i g h t estimates. T h i s i s considered t o be reason- a b l e agreement. It shows t h e wind t u n n e l d r a g t o be h i g h e r than f l i g h t d r a g as would be expected, b u t a t approximately t h e same l e v e l . An increment of about 20 c o u n t s would be expected f o r t h e d i f f e r e n c e i n Reynolds numbers between t h e two cases.

The e f f e c t s of adding the PTA n a c e l l e t o t h e G I 1 a i r c r a f t are shown i n Figure 2 9 . A t z e r o l i f t , t h e impact of t h e n a c e l l e i s t o add about 30 c o u n t s of d r a g , but i n t h e range of l i f t c o e f f i c i e n t s from 0.1 t o 0.3, t h e n a c e l l e adds about 20 t o 25 counts of drag.

F i g u r e 30 shows t h e a d d i t i o n a l d r a g increment due t o the a d d i t i o n of t h e wing t i p booms. Over t h e range of moderate l i f t c o e f f i c i e n t s , t h e booms add about 20 c o u n t s of drag. The a c o u s t i c boom was n o t added t o t h e low-speed model because r e s u l t s from t h e high-speed t u n n e l tests had a l r e a d y shown t h e d r a g of t h i s a d d i t i o n t o be n e g l i g i b l e . The t o t a l increment due t o PTA m o d i f i c a t i o n s from t h e low-speed wind t u n n e l tests i s approximately 40 t o 4 5 c o u n t s of drag.

Because Reynolds numbers were h i g h e r i n t h e high-speed wind t u n n e l tests, i t would be expected t h a t d r a g increments from t h e high-speed tests would be s m a l l e r than t h o s e from t h e low-speed tests. T h i s can be s e e n i n Figure 31 where d r a g increments f o r t h e PTA b u i l d u p are p l o t t e d a g a i n s t Mach number. These d a t a show a d r a g c r e e p e f f e c t a t Mach numbers above 0.6 and a s h a r p t r a n s o n i c d r a g rise above Mach 0.7.

The e f f e c t of Mach number on d r a g p o l a r s f o r t h e PTA model i s shown i n F i g u r e 32. I n t h i s f i g u r e , t h e i n f l u e n c e of l i f t c o e f f i c i e n t on t h e o n s e t of t r a n s o n i c d r a g rise can be seen.

I n t h e PTA Program, p r o v i s i o n s were made f o r v a r y i n g n a c e l l e i n c i d e n c e i n o r d e r t h a t i n f l o w a n g l e t o t h e propfan could be v a r i e d over a wide range. The e f f e c t s of n a c e l l e i n c i d e n c e on d r a g are shown i n Figure 3 3 . Drag a p p e a r s t o be u n a f f e c t e d by changes i n n a c e l l e i n c i d e n c e from - 1 t o +2 d e g r e e s , but i n c r e a s e s s i g n i f i c a n t l y when n a c e l l e tilt is changed t o -3 degrees.

The e f f e c t s of f l a p s on PTA model d r a g a r e shown i n F i g u r e 34 where d r a g p o l a r s are shown f o r f l a p a n g l e s of 0 , 20, and 4 0 d e g r e e s , and f o r both t h e G I 1 and t h e PTA models. Drag performance w i t h f l a p s is n o t g r e a t l y a f f e c t e d by t h e a d d i t i o n of t h e PTA m o d i f i c a t i o n s .

7.2 PRESSURE DISTRIBUTIONS The flow f i e l d around a p r o p e l l e r n a c e l l e i n s t a l l e d on a swept wing i s q u i t e complex. T h e r e f o r e , t h e PTA model w a s instrumented w i t h a number of p r e s s u r e p o r t s on t h e n a c e l l e and on t h e wing n e a r t h e n a c e l l e , so t h a t t h e flow f i e l d s could be i n t e r p r e t e d and t h e a n a l y t i c a l t o o l s used i n d e s i g n could be e v a l u a t e d .

Figure 35 shows a comparison of p r e d i c t e d and measured p r e s s u r e s on t h e n a c e l l e . The wire diagrams a t t h e t o p of t h e f i g u r e show the a n a l y t i c a l model, and t h e shaded areas show t h e l o c a t i o n of p a n e l s where p r e s s u r e was computed. For t h i s p r o p e l l e r - o f f , Mach 0.4 c o n d i t i o n , i t can be seen that t h e a n a l y t i c a l method d i d an e x c e l l e n t j o b of p r e d i c t i n g e x p e r i m e n t a l r e s u l t s .

F i g u r e 36 shows a comparison of prop-on wind t u n n e l d a t a w i t h pre- d i c t e d v a l u e s f o r p r e s s u r e s on t h e wing j u s t inboard and j u s t outboard of t h e n a c e l l e . On t h e inboard s i d e , t h e a d d i t i o n a l upwash caused by t h e Power-Off Drag of PTA Nacelle Figure 29.

e -

O P T A WITHOUT BOOMS O P T A WITH BOOMS

f

l 0 0 . 1 2 Figure 30. Drag of Tip Booms M A C H N U M B E R Figure 31.

Drag of PTA Components a t High Speed

t

i -0.2 J DRAG COEFFICIENT, Cg Figure 32.

Effect of Mach Number on PTA Drag Polars I - 1 . 0 - - 0 . 8

-

- 0.6 cL - 0.4

-

- 0 . 2 - -0.02 - - - 0 . 2

-

I

--0.4 Figure 33. Effect of Varying Nacelle Incidence on Drag

e -

Figure 34. Effects of Flap Deflection on Drag

- -

PTA GULFSTREAM I1 F . a OUADPAN DATA M A C I ~ 0.4. ALPIIA = 2 . 0 0 PTA GULFSTREAM II MACH = -40 ALPHA = 2.0 PROP OFF

- QUAOPAN ,ANALY SIS OWIND-TUNNEL,DATA

OUTBOARD SIDE INBOARD SIDE - 1 . 2 T O P OF COWL - 1 . 2 OF NACELLE OF NACELLE CP CP 0.Q4 0 4 0 4 - 0 00 0 1 S 0 50 0 75 1.00 0 00 0 2 5 0 50 0 15 1 00 0 00 0.15 0 50 0 75 I 00 X I L XI. X , L

Figure 3 5 . Nacelle Surface Pressure Distributions - Prop-Off

q = 1.0 ' O WIND-TUNNEL D A T A

- QUADPAN ANALYSIS

2 WING S T A T I O N 1 WING S T A T I O N rl= 0.290 3 NACEELE INBOARD S T A T I O N 9= 0.320 -1.1- CP 0 1 4 0 1 4 *.os 0.15 0.50 0 75 1 00 *.os 0.15 0.50 0 75 1 00 x 'C x 'C 4 T O P OF NACELLE 5 OUTBOARD OF NACELLE 6 WING STATION, q = 0.480 6 WING STATION, q = 0.480 0 0 n n

- -

- 1 . 1 0 0 4 4 CP 0 0 f f -0.1 CP CP

s s

0 0 .oI .oI : : 0.1

Figure 36. Wing Surface Pressure Distributions - Mach 0.4

s l i p s t r e a m s w i r l g i v e s rise t o p r e s s u r e s u c t i o n peaks i n t h e l e a d i n g edge r e g i o n of the wing. On t h e outboard s i d e t h e downwash on t h e lower s u r f a c e g i v e s rise t o s u c t i o n peaks i n t h e lower s u r f a c e l e a d i n g edge r e g i o n . P r e d i c t i o n s a t Mach 0 . 4 a r e i n e x c e l l e n t agreement w i t h t h e test data. On t h e inboard s i d e , t h e magnitude of t h e s u c t i o n peak i s accu- r a t e l y predicted. The inadequacies of QUADPAN paneling a t t h e t r a i l i n g edge w i t h vanishing s t r i p s , however, g i v e s l o c a l i z e d erroneous p r e s s u r e c o e f f i c i e n t s .

Outboard of t h e n a c e l l e , c o r r e l a t i o n of t h e o r y and experimental d a t a is e q u a l l y good. T h i s c o r r e l a t i o n shows t h a t t h e QUADPAN a n a l y s i s pre- d i c t s q u i t e a c c u r a t e l y t h e f l o w f i e l d a t subsonic Mach numbers on t h e complex geometry of t h e i n t e g r a t e d wing/nacelle combination. I n a d d i t i o n , t h e agreement i n d i c a t e s t h a t t h e p r e d i c t e d slipstream has t h e c o r r e c t induced v e l o c i t y and s w i r l d i s t r i b u t i o n f o r t h e o p e r a t i n g advance r a t i o and b l a d e angle s e t t i n g .

A t t h e higher Mach number of 0.7, Figure 37 shows t h a t t h e v e l o c i t y i n t h e channel between t h e n a c e l l e and t h e f u s e l a g e a c c e l e r a t e s t o super- s o n i c v a l u e s with t h e p r e s s u r e peak moving a f t and t h e n d e c e l e r a t e s through a weak shock wave. QUADPAN p r e d i c t s t h e s u c t i o n peak, b u t cannot p r e d i c t t h e supersonic flow behind t h e peak.

F i g u r e s 38 and 39 show t h a t t h e a n a l y s i s program was e f f e c t i v e f o r p r e d i c t i n g prop-on flow around t h e n a c e l l e a t Mach numbers of 0.7 and 0.8, r e s p e c t i v e l y . Again, however, s i n c e QUADPAN i s n o t a t r a n s o n i c code, t h e flow p r e d i c t i o n on t h e inboard s i d e of t h e n a c e l l e i s not so good where t h e s w i r l i n g s l i p s t r e a m i n c r e a s e s l o c a l angle of a t t a c k and c r e a t e s l o c a l r e g i o n s of supersonic flow.

F i g u r e 40 shows wing s u r f a c e p r e s s u r e s j u s t inboard and j u s t outboard of t h e PTA n a c e l l e l o c a t i o n f o r t h e GI1 and PTA models. On t h e inboard s i d e of t h e n a c e l l e , t h e e f f e c t of t h e n a c e l l e causes a l o c a l a c c e l e r a t i o n n e a r t h e wing l e a d i n g edge. There i s , however, no evidence of any s i g n i f - i c a n t flow s e p a r a t i o n . O n t h e outboard s i d e of t h e n a c e l l e , t h e primary e f f e c t of the n a c e l l e appears t o be a lowering of t h e l o c a l a n g l e of a t t a c k and a s l i g h t r e d u c t i o n i n l i f t .

S i m i l a r d a t a a r e shown i n Figure 41 for a Mach number of 0 . 8 . Here i t can be seen t h a t some shock waves a r e formed on both inboard and out- board s i d e s of t h e n a c e l l e . The h i g h e r l o c a l Mach numbers on t h e inboard s i d e of t h e n a c e l l e r e s u l t i n p a r t from t h e unsweeping e f f e c t of t h e n a c e l l e on t h e wing and i n p a r t from t h e a c c e l e r a t i o n i n t h e channel formed by t h e n a c e l l e , wing, and f u s e l a g e . These e f f e c t s probably account f o r most of t h e PTA d r a g increment a t t h e h i g h e r Mach numbers.

The e f f e c t s of angle-of-attack change on t h e PTA wing chordwise pres- s u r e d i s t r i b u t i o n s a t Mach 0 . 8 a r e shown i n Figure 42. A s is normal w i t h a n i n c r e a s e i n a n g l e of a t t a c k , flow v e l o c i t i e s a r e i n c r e a s e d over t h e wing and decreased under t h e wing--resulting i n i n c r e a s e d c i r c u l a t i o n It would be expected t h a t an i n c r e a s e i n d r a g would accompany t h i s l i f t .

a t low a n g l e s of a t t a c k b u t i n c r e a s i n g r a p i d l y as l i f t increase--moderate a n g l e of a t t a c k i n c r e a s e s . C a l c u l a t i o n s confirm t h a t , i n t h e a n g l e of a t t a c k range from z e r o t o about t h r e e degrees, t h e l i f t - t o - d r a g r a t i o was approximately 10, b u t dropped t o 8.5 a t angle of a t t a c k of 4 degrees.

The e f f e c t of i n c r e a s i n g Mach number is shown i n Figure 4 3 . B e l o w t h e c r i t i c a l f r e e s t r e a m Mach number, t h e s u c t i o n peak i s c l o s e t o t h e wing l e a d i n g edge and i n c r e a s e s with Mach number. I n t h i s range, t h e flow M=0.7, d = 2 -1.60 QUADPAN WIND TUNNEL 1) =.472 - 1 . 2 0 -1.20 a V +' -0.80 z w V

-

5 -0.40 W W 2 -0.00 v) v) W 0.40 0.80 1 . 2 0

Figure 3 7 . Wing Surface Pressure Distributions - Mach 0 . 7

PTA GULFSTREAM I f MACH 0.7. ALPHA = 2O PROP ON _ -

TOP \ fl _--

/-,#

- QUADPAN ANALYSIS

0 WIND-TUNNEL DATA OUTBOARD SIDE INBOARD SIDE

- 1 . 1 1 TOP OF COWL

OF NACELLE OF NACELLE

c~~~ I

- 0 0 -0.10 cp - 6

'1

I Figure 38. Nacelle Surface Pressure Distributions - Prop-On, Mach 0 . 7 a

e -

PTA GULFSTREAM ll MACH = 0.80 ALPHA = 2 ' PROP ON

- QUADPAN ANALYSIS

0 WIND TUNNEL D A T A

OUTBOARD SIDE I lNBOARD SIDE

-1 .I TOP OF COWL O F NACELLE -0.Q i ' 0.4 0 00 0.13 0.50 1.15 1.00 0.00 0 . 2 5 0.50 0.75 1.00 < 0 . 0 0 0.25 0 50 0.75 I 00 * i X I L - X I L X I L Figure 3 9 . Nacelle Surface Pressure Distributions - Prop-On, Mach 0.8 MODEL PTA Figure 40. Effect of PTA Nacelle on Wing Section Pressures- Mach 0.7 MODEL A PTA - 1 . 6 0 1 A q =.328 a U I-' -0.80 z w

u

u . -0.90 U W V I I w -0.00 ' 0.20 0.40 0.60 0.8 K

r- FRACTION OF CHORD. X I C '

m m m m W W g 0.40 0.40 a 0.80

o'80L 1 . 2 0

1 . 2 0 Figure 41. Effect o f PTA Nacelle on Wing Section Pressures- Mach 0.8 CI M-0 .8 U U Effect o f Angle o f Attack on Wing Section Pressures Figure 4 2 .

t ) = . 3 2 8

- - - 0.699

- - - 0.798

-- 0 . 8 5 2

0 .

U ff = 2 0 -1.20 I-- z W

-

v

-0.80 u .

L W a U U -0.40 W a m m w -0.00 cz P 0.40

0.80 J

Effect of Mach Number on Wing Section Pressures Figure 43.

behind t h e peak recompresses g r a d u a l l y w i t h no s i g n s of flow s e p a r a t i o n .

Above t h e c r i t i c a l Mach number, however, t h e s u c t i o n peak moves a f t and t h e flow recompresses through a shock wave t h a t becomes s t r o n g e r w i t h i n c r e a s i n g Mach number. I n t h i s range, t h e r e i s danger of flow s e p a r a t i o n due t o t h e s t r o n g i n t e r a c t i o n between t h e shock wave and t h e boundary l a y e r . A t Mach 0.8, t h e r e i s no evidence of e x t e n s i v e s e p a r a t i o n , a l t h o u g h t h e r e may be some l o c a l l y a t t h e f o o t of t h e shock. A t Mach 0.85, however, t h e r e i s evidence of s e p a r a t i o n at t h e t r a i l i n g edge that c a u s e s t h e shock wave t o move forward. T h i s is e v i d e n t i n t h e p r e s s u r e d i s t r i b u t i o n s on t h e outboard s i d e of t h e n a c e l l e where s e p a r a t i o n f i r s t occurs. The wing lower s u r f a c e p r e s s u r e s a l s o d e c r e a s e w i t h i n c r e a s i n g Mach number, and a t Mach 0.85, s o n i c v e l o c i t i e s are reached outboard of t h e n a c e l l e and t e r m i n a t e through a shock wave. T h i s t r a n s o n i c behavior i s r e s p o n s i b l e f o r t h e d r a g rise t h a t starts a t about Mach number 0.7.

7 . 3 STABILITY AND CONTROL I n t h i s s e c t i o n , aerodynamics d a t a from t h e wind t u n n e l tests are p r e s e n t e d and d i s c u s s e d from t h e s t a n d p o i n t of t h e impact of t h e s e d a t a on s t a b i l i t y and c o n t r o l of t h e PTA c o n f i g u r a t i o n . It was intended t h a t wind t u n n e l d a t a would be used i n t h e PTA a i r c r a f t d e s i g n by adding t h e wind- tunnel-derived f o r c e and moment increments t o f u l l - s c a l e G I 1 d a t a s u p p l i e d by Gulfstream Aerospace Corporation. I n t h e a n a l y s i s and d i s c u s s i o n , t h e r e f o r e , major emphasis is placed on increments between G I 1 and PTA models. It should be remembered t h a t t h e PTA wind t u n n e l model w a s a m i r r o r image of t h e PTA a i r c r a f t - - w i t h t h e n a c e l l e on t h e right-hand wing i n t h e wind t u n n e l i n s t e a d of t h e l e f t .

7.3.1 L i f t and P i t c h - Low Speed

The low-speed l i f t and p i t c h i n g moment c o e f f i c i e n t s f o r t h e PTA model i n t h e c l e a n , t a k e o f f , and l a n d i n g c o n f i g u r a t i o n s , i n t h e "ferry" mode, and with a f e a t h e r e d propfan, are compared w i t h t h e G I 1 i n Figure 4 4 .

I n t h e c l e a n c o n f i g u r a t i o n , t h e major changes due t o t h e PTA i n s t a l - l a t i o n were an almost c o n s t a n t i n c r e a s e i n l i f t c o e f f i c i e n t , over t h e o p e r a t i o n a l range of a n g l e of a t t a c k , and a small r e d u c t i o n i n maximum l i f t c o e f f i c i e n t a t a lower angle of a t t a c k . Very l i t t l e e f f e c t i s shown due t o t h e a d d i t i o n of a f e a t h e r e d p r o p e l l e r . The p i t c h i n g moment d a t a show no e f f e c t on CMo but a s i g n i f i c a n t d e s t a b i l i z i n g e f f e c t on dCM/dCL t h a t i s f u r t h e r i n c r e a s e d by t h e f e a t h e r e d p r o p e l l e r . These e f f e c t s are a r e s u l t of l i f t generated on t h e n a c e l l e forward of t h e c e n t e r of g r a v i t y and t h e r e s u l t a n t changes i n wing l i f t due t o t h e n a c e l l e and t i p booms.

I n t h e t a k e o f f f l a p c o n f i g u r a t i o n , a similar l i f t i n c r e a s e occurred due t o t h e PTA i n s t a l l a t i o n and a similar l o s s of maximum l i f t c o e f f i c i e n t a t a lower a n g l e of a t t a c k . The f e a t h e r e d p r o p e l l e r reduced t h e l i f t increment i n t h e o p e r a t i o n a l a n g l e of a t t a c k range and a l s o f u r t h e r reduced t h e maximum l i f t c o e f f i c i e n t . T h i s r e d u c t i o n i n l i f t d e c r e a s e s t h e s t a b i l i t y l e v e l w i t h a f e a t h e r e d p r o p e l l e r a t t h e h i g h e r l i f t c o e f f i - c i e n t s . A small d e s t a b i l i z i n g change i n t a i l l i f t due t o t h e n a c e l l e c o n t r i b u t e s t o t h i s e f f e c t , but t h e major changes occur on t h e wing and n a c e l l e .

CONFIG PROP FLAPS ---- 0 GI1 0 DEC 0 P T A OFF 0 DEG A PTA FEATH ----

+ GI1

X PTA OFF FEATH 0 PTA ----

+ GI1

X PTA OFF Z PTA FEATH Figure 44.

Effects of PTA Modifications on Lift and Pitching Moment I n t h e landing c o n f i g u r a t i o n , t h e l i f t i n c r e a s e due t o t h e PTA i n s t a l l a t i o n remained t h e same as t h e t a k e o f f mode, and t h e e f f e c t of adding f e a t h e r e d p r o p e l l e r s was a l s o t h e same. The maximum l i f t l o s s due E O t h e n a c e l l e and booms was reduced, and s t a l l occurred a t a s l i g h t l y lower angle of a t t a c k . The s t a b i l i t y d e c r e a s e r e l a t i v e t o t h e G I 1 is v e r y small, and t h e e f f e c t of f e a t h e r e d p r o p e l l e r s is less d e s t a b i l i z i n g than t h e t a k e o f f f l a p l e v e l .

The reduced l e v e l of s t a t i c s t a b i l i t y f o r t h e PTA c o n f i g u r a t i o n a t a l l f l a p s e t t i n g s i s almost e x a c t l y balanced by a forward s h i f t i n t h e c e n t e r of g r a v i t y envelope r e l a t i v e t o t h e GII. Thus, t h e a c t u a l minimum s t a t i c s t a b i l i t y margin remains t h e same as t h e G I 1 a t t h e d e s i g n a f t c e n t e r of g r a v i t y . The r e d u c t i o n i n maximum l i f t c o e f f i c i e n t r e s u l t s i n a s l i g h t i n c r e a s e i n minimum o p e r a t i o n a l speeds r e l a t i v e t o t h e G I I . These i n c r e a s e s are of l i t t l e consequence.

The e f f e c t of propfan power on l i f t and p i t c h i n g moment, f l a p s up, i s shown i n F i g u r e 45. The l i f t increment due t o power i n c r e a s e d w i t h a n g l e of a t t a c k and t h e maximum l i f t c o e f f i c i e n t i n c r e a s e d . The angle of a t t a c k f o r maximum l i f t d i d n o t change r e l a t i v e t o t h e prop-off value. T h i s e f f e c t i s t h e r e s u l t of p r o p e l l e r normal f o r c e and t h e s l i p s t r e a m e f f e c t on t h e n a c e l l e and wing. The incremental e f f e c t of power on l i f t would be less than shown h e r e f o r f u l l - s c a l e a i r c r a f t due t o t h e change i n Reynolds number, b u t t h e same t r e n d s w i l l p r e v a i l . P i t c h i n g moment becomes more p o s i t i v e with power, and t h e s t a b i l i t y l e v e l d e c r e a s e s : t h e s e l e v e l s w i l l be s l i g h t l y less a t f u l l - s c a l e Reynolds number.

L a t e r a l - D i r e c t i o n a l - Low Speed - Zero S i d e s l i p

7.3.2 The low-speed s i d e f o r c e v a r i a t i o n with angle of a t t a c k , a t z e r o s i d e s l i p , is shown i n F i g u r e 46 f o r t h e GI1 and t h e PTA c o n f i g u r a t i o n s , f l a p s up. Small s i d e f o r c e s are a p p a r e n t f o r t h e b a s i c G I 1 and can be a t t r i b u t e d t o t h e e f f e c t s of t h e support system and t u n n e l asymmetries.

The a d d i t i o n of t h e n a c e l l e and booms c r e a t e d a n e g a t i v e s i d e f o r c e incre- ment a t a n g l e s of a t t a c k above zero. This increment has c o n t r i b u t i o n s from t h e n a c e l l e , t h e booms, and from flow changes a t t h e f i n . The a d d i t i o n of t h e f e a t h e r e d propfan r e s u l t e d i n an i n c r e a s e i n p o s i t i v e s i d e f o r c e t h a t i s due t o wing-body changes and a small f l o w f i e l d change a t t h e f i n .

Yawing moment v a r i a t i o n w i t h a n g l e s of a t t a c k , a t z e r o s i d e s l i p , w a s e s s e n t i a l l y z e r o f o r a l l o p e r a t i o n a l a n g l e s of a t t a c k . The yawing moment due t o n a c e l l e d r a g was o f f s e t by t h e yawing moment due t o induced s i d e l o a d s on t h e f i n . The h i g h a n g l e of a t t a c k yawing moments are a r e s u l t of low Reynolds number wing s e p a r a t i o n and would s h i f t t o h i g h e r a n g l e s a t f u l l - s c a l e c o n d i t i o n s .

The r o l l i n g moment e f f e c t i s l a r g e l y due t o t h e l i f t on t h e n a c e l l e and t h e l i f t induced on t h e wing. This r o l l i n g moment is e s s e n t i a l l y c o n s t a n t with a n g l e of a t t a c k and could be e a s i l y balanced by a i l e r o n d e f l e c t i o n f o r trimmed f l i g h t . On t h e PTA, t h e n a c e l l e weight is n o t completely balanced, and t h e a i r c r a f t c e n t e r of g r a v i t y i s d i s p l a c e d t o t h e same s i d e as t h e n a c e l l e , c r e a t i n g a r o l l i n g moment i n t h e o p p o s i t e d i r e c t i o n . T h i s mass imbalance t e n d s t o c o u n t e r t h e aerodynamic e f f e c t s o t h a t t h e a i r c r a f t i s self-trimming a t i n t e r m e d i a t e speeds.

1 TC 0 PROP OFF 0 1.8 0.1 A 1.0 0 . 6 4- 0.85 0.9 FLAPS U P Figure 4 5 . Effects of Propfan Power on Aerodynamic Characteristics

in Pitch - Flaps U p

CONFIC PROP FLAPS ---- 0 OEG 0 C l i 0 PTA OFF 0 DEG A PTA FEATH 0 DEC z V ANGLE O F ATTACK, a ( D E C ) Effects o f PTA Modifications on Side Force, Yawing Figure 46.

Moment, and Rolling Moment - Flaps Up

With f l a p s down i n t h e t a k e o f f and l a n d i n g modes, t h e e f f e c t of t h e n a c e l l e r e l a t i v e t o t h e GI1 is v e r y small as shown i n Figures 47 and 48.

There are e f f e c t s i n s i d e f o r c e and r o l l i n g moment f o r t h e GI1 t h a t are t h e r e s u l t of an u n i d e n t i f i e d asymmetry i n t h e f l a p c o n f i g u r a t i o n t h a t should n o t unduly i n f l u e n c e t h e incremental e f f e c t of t h e PTA i n s t a l l a - t i o n .

The e f f e c t of power on s i d e f o r c e , yaw, and r o l l i n g moment as a f u n c t i o n of angle of attack a t z e r o s i d e s l i p is shown i n Figure 49. The small o f f s e t i n s i d e f o r c e w i t h p r o p e l l e r s o f f r o t a t e d about an a n g l e of a t t a c k of 4 degrees a s power increased. Normal c r u i s e a n g l e of a t t a c k is around 3 t o 4 d e g r e e s a t t h e lower speeds a s s o c i a t e d w i t h h i g h power c o e f f i c i e n t s ; hence, t h e change i n bank a n g l e r e q u i r e d t o balance t h e s i d e f o r c e a t z e r o s i d e s l i p would remain q u i t e small. The d a t a show an i n c r e a s e i n s i d e f o r c e i n t h e n e g a t i v e d i r e c t i o n w i t h power i n c r e a s e t h a t f l a t t e n s t o a c o n s t a n t v a l u e a t h i g h TC.

The yawing moment change w i t h power f o l l o w s t h e i n c r e a s e i n t h r u s t very c l o s e l y w i t h l i t t l e n e t aerodynamic i n f l u e n c e . The l e v e l of yawing moment c o e f f i c i e n t shown h e r e a t a T of 0.9 i s v e r y c l o s e t o t h e maximum yawing moment a v a i l a b l e from t h e $udder and would t h e r e f o r e provide a l i m i t on t h e minimum f l i g h t speed w i t h f u l l propfan power. However, t h e r o l l i n g moment due t o power reached a l i m i t i n g c o n d i t i o n a t a TC of 0.6.

The r o l l i n g moment d a t a shown i n F i g u r e 49 is made up of p r o p e l l e r normal f o r c e , s i d e f o r c e , propeller-slipstream-induced l i f t on t h e wing, and induced loads on t h e f i n and h o r i z o n t a l t a i l . The l e v e l i n c r e a s e d w i t h both angle of a t t a c k and TC and is t h e r e f o r e c r i t i c a l t o low-speed f l i g h t . For t h e PTA a i r c r a f t , t h e l i m i t i n g c o n d i t i o n f o r low-speed f l i g h t i s based on t h e t r i m c o n d i t i o n of no more t h a n 50-percent wheel throw t o balance t h e o f f s e t roll.

7 . 3 . 3 L a t e r a l - D i r e c t i o n a l - Low Speed - V a r i a b l e S i d e s l i p The v a r i a t i o n of s i d e f o r c e , yawing moment, and r o l l i n g moment w i t h s i d e s l i p a n g l e , shown i n Figure 50, demonstrates t h e small i n f l u e n c e of t h e n a c e l l e and booms on a i r c r a f t l a t e r a l - d i r e c t i o n a l s t a b i l i t y . R e l a t i v e t o t h e G I I , t h e s i d e f o r c e v a r i a t i o n w i t h s i d e s l i p i n c r e a s e d , and t h e yawing moment and r o l l i n g moment v a r i a t i o n decreased. These e f f e c t s are p r i m a r i l y due t o t h e added n a c e l l e s i d e area w i t h l i t t l e o r no i n f l u e n c e from t h e v e r t i c a l f i n . The n o n - l i n e a r i t y i n t h e r o l l i n g moment c u r v e s I s a normal e f f e c t of wing load d i s t r i b u t i o n changes w i t h s i d e s l i p t h a t d e c r e a s e w i t h i n c r e a s i n g Reynolds number f o r t h e f u l l - s c a l e a i r c r a f t . The o f f s e t i n r o l l a t z e r o s i d e s l i p i s t h e v a l u e f o r 3-degrees a n g l e of a t t a c k and i s d i s c u s s e d i n t h e previous s e c t i o n .

The e f f e c t of power on t h e low-speed, l a t e r a l - d i r e c t i o n a l d a t a is shown i n Figure 51. s i d e f o r c e due t o s i d e s l i p (dCy/dB) was i n c r e a s e d The by propfan power due t o s l i p s t r e a m e f f e c t s on t h e wing and n a c e l l e .

Yawing moment due t o s i d e s l i p (dCn/dB) w a s s l i g h t l y decreased by propfan power p r i m a r i l y due t o change i n t h e wing l o a d i n g and d i r e c t p r o p e l l e r e f f e c t s . R o l l i n g moment due t o s i d e s l i p w a s reduced s l i g h t l y w i t h propfan power as a r e s u l t of wing l o a d changes. The o f f s e t s a t z e r o s i d e s l i p i n s i d e f o r c e , yawing moment, and r o l l i n g moment are a p p r o p r i a t e f o r 3 d e g r e e s of a n g l e of a t t a c k and are d i s c u s s e d i n t h e p r e v i o u s s e c t i o n .

CONFIC PROP FLAPS U C l l _--- 20 DEC O P T A OFF 20 DEC APTA FEATIi 20 DEC Figure 4 7 . Effects of PTA Modifications on Side Force, Yawing Moment, and Rolling Moment - Flaps 20-Degrees

*-

CONFIG PROP FLAPS ___- 0 GI1 40 DEG 0 PTA OFF 40 DEG A PTA FEATH 40 DEC x U Figure 48. Effects of PTA Modifications on Side Force, Yawing Moment, and Rolling Moment - Flaps 40-Degrees J TC 0 PROP OFF 0 1 . 8 0.1 A 1.0 0 . 6 -k 0 . 8 5 0.9 FLAPS U P BETA = 0 DEG Figure 49. Effects o f Power on Aerodynamic Characteristics in

Pitch - Flaps Up

CONFIGURATION 0 GI1 0 PTA (PROP OFF) A PTA (PROP FEATHERED) FLAPS UP ALPHA = 3 DEC Is V.7 V r-' Z w u .

LL w V + t - Z w I E v f -I -1 (L Figure 5 0 . Effects of PTA Modifications on Sideslip Characteristics- Flaps Up

*-

J TC c] PROP OFF FLAPS UP 0 1.8 0 . 1 0 1.0 0.6 ALPtlA = 3 DEG -I. 0.85 0.9 C u e I - Z Z W W I I I I V f -I -I K Figure 5 1 . Effects of Propfan Power on Aerodynamic Characteristics

in Sideslip - Flaps Up

S i m i l a r d a t a are shown i n F i g u r e s 52 and 53 f o r t h e t a k e o f f and land- i n g c o n f i g u r a t i o n s , r e s p e c t i v e l y . The same l e v e l of i n c r e a s e i n s i d e f o r c e v a r i a t i o n and d e c r e a s e i n yawing moment and r o l l i n g moment w i t h s i d e s l i p a n g l e occurred r e l a t i v e t o t h e G I 1 as f o r t h e flaps-up c o n f i g u r a t i o n . The o f f s e t i n r o l l i n g moment f o r t h e b a s i c G I 1 i s a g a i n apparent.

No propfan powered d a t a were o b t a i n e d i n t h e wind t u n n e l f o r t h e f l a p s - d e f l e c t e d cases because t h e PTA a i r c r a f t was not designed t o o p e r a t e under t h o s e c o n d i t i o n s .

7 . 3 . 4

C o n t r o l Effectiveness - Low Speed

E l e v a t o r e f f e c t i v e n e s s f o r t h e t h r e e f l a p c o n f i g u r a t i o n s i s shown i n F i g u r e 54. With t h e f l a p s up, t h e r e was s l i g h t l y less p i t c h i n g moment due t o e l e v a t o r d e f l e c t i o n i n t h e PTA c o n f i g u r a t i o n r e l a t i v e t o t h e G Z I . Both show f u r t h e r l o s s e s a t n e g a t i v e a n g l e s of a t t a c k . These d i f f e r e n c e s are l a r g e l y due t o t h e v a r i a t i o n of t a i l angle of a t t a c k w i t h f l a p s e t t i n g and a t n e g a t i v e v a l u e s ( f u s e l a g e a n g l e of a t t a c k less t h a n 1.4 show t h a t d e g r e e s ) t h e e f f e c t i v e n e s s of full-up e l e v a t o r d e f l e c t i o n diminished b u t was n o t s t a l l e d . For 20 d e g r e e s of f l a p , t h e n e g a t i v e t a i l angle of a t a f u s e l a g e angle of a t t a c k of 8 d e g r e e s and showed a t t a c k range began t h e t y p i c a l d e c r e a s e i n e f f e c t i v e n e s s through zero. For f l a p s a t 48 d e g r e e s , t h e t a i l angle of a t t a c k i s always n e g a t i v e , and t h e secondary "plateau" has a l r e a d y been reached. A t z e r o f u s e l a g e a n g l e of a t t a c k , t a i l s t a l l began w i t h f u l l up e l e v a t o r , t h u s t h e s h a r p d e c r e a s e i n e l e v a t o r e f f e c t i v e n e s s . Under t h e s e c o n d i t i o n , t h e t o t a l t a i l l i f t is a lmos t cons t a n t .

A l l t h e s e e f f e c t s are h i g h l y Reynolds number dependent, and a t f u l l - s c a l e c o n d i t i o n s , i t i s a n t i c i p a t e d t h a t no l o s s of e l e v a t o r e f f e c t i v e n e s s , e i t h e r on t h e G I 1 or t h e PTA, would a c t u a l l y occur. The t r a i l i n g edge down e l e v a t o r of 10 degrees shows l i t t l e o r no e f f e c t of a n g l e of a t t a c k on PTA c o n f i g u r a t i o n .

are shown i n F i g u r e s 55 through 5 7 f o r t h e Rudder e f f e c t i v e n e s s d a t a G I 1 and PTA models i n t h e c l e a n , t a k e o f f , and l a n d i n g c o n f i g u r a t i o n s as a f u n c t i o n of f u s e l a g e a n g l e of a t t a c k . Very l i t t l e d i f f e r e n c e i n effec- t i v e n e s s occurred f o r any of t h e axes r e l a t i v e t o t h e G I I , b u t t h e r e does appear t o be a s l i g h t i n c r e a s e due t o propfan power, f l a p s up.

S p o i l e r e f f e c t i v e n e s s d a t a are shown i n Figure 58 as increments due t o s p o i l e r d e f l e c t i o n about a l l axes f o r t h e G I 1 and t h e PTA configura- t i o n s . It should be noted t h a t t h e s i n g l e outboard s p o i l e r p a n e l i s used f o r both GI1 and PTA d a t a and is mounted on t h e r i g h t wing w i t h t h e PTA n a c e l l e . The l i f t decrement due t o s p o i l e r d e f l e c t i o n i s i n c r e a s e d by t h e presence of t h e propfan n a c e l l e . The n a c e l l e i n c r e a s e d t h e o u t e r wing span l o a d i n g , t h u s c r e a t i n g more l i f t f o r t h e s p o i l e r t o reduce. The a d d i t i o n of a f e a t h e r e d propfan reduced t h e l i f t decrement due t o t h e s p o i l e r back t o t h e same l e v e l as t h e G I 1 c o n f i g u r a t i o n f o r a n g l e s of a t t a c k up t o 6 d e g r e e s and t h e n shows f u r t h e r d e g r a d a t i o n a t h i g h e r a n g l e s as t h e l i f t due t o t h e n a c e l l e increased. T h i s occurred because t h e f e a t h e r e d p r o p e l l e r i n f l u e n c e d t h e o u t e r wing and reduced t h e l i f t g a i n s from t h e n a c e l l e . The p i t c h i n g moment increment due t o t h e s p o i l e r s i s i n t h e a i r c r a f t nose-up d i r e c t i o n and is c o n s i s t e n t w i t h t h e l i f t r e d u c t i o n behind t h e moment c e n t e r . S i d e f o r c e changes are n e g l i g i b l e , and a small yawing moment increment due t o s p o i l e r d r a g is apparent. The r o l l due t o

*-

CONFIGURATION FLAPS 20 DEC a GI1 0 PTA (PROP OFF) ALPHA = 3 DEG A PTA (PROP FEATHERED) Figure 5 2 . Effects of PTA Modifications on Sideslip Characteristics- Flaps 20-Degrees CONFIGURATION FLAPS Y O DEG 0 GI1 0 PTA (PROP OFF) ALPHA = 3 DEG A PTA (PROP FEATHERED) + PTA (PROP WINDMILLING)

c

z

w

v

U .

U W V I- Z ' W z z u f -1 -t K Figure 5 3 . Effects of PTA Modifications on Sideslip Characteristics- Flaps 40-Degrees CONFIGURATION ELEV 0 G I 1 - 2 5 0 ,I +10 A PTA (TC=O.I) -25 -6 I, I, +IO P T A (TC-0.9) -25 +10 Figure 5 4 . Elevator Effectiveness CONFIGURATION RUDDER 0 GI1 - 2 5 FLAPS UP 0 PTA ( T C = O . l l - 2 5 BETA = 0 A PTA (TC=0.9) -25

Rudder Effectiveness in Pitch - Flaps Up

Figure 5 5 .

CONFIGURATION RUDDER 0 GI1 -2s FLAPS 2 0 0 PTA (FEATHERED) - 2 5 BETA = 0 !.

Rudder Effectiveness in Pitch - Flaps 20-Degrees

Figure 5 6 .

*-

CONFIC'N RUDDER 0 GII - 2 5 FLAPS QO 0 PTA [FEATHERED) -25 BETA = 0

Figure 57. Rudder Effectiveness in Pitch - Flaps 40-Degrees

CONFIGURATION R T SPLR 0 G I I -35 0 PTA (PROP OFF) -35 n PTA (FEATHERED) -35 FLAPS UP BETA = 0 LE OF ATTACK, cy IDEG))

Figure 58. Spoiler Effectiveness in Pitch - Flaps U p

s p o i l e r d e f l e c t i o n had t h e same shape as t h e l i f t r e d u c t i o n and remained almost c o n s t a n t t o h i g h a n g l e s of a t t a c k . F u l l - s c a l e , t h e s e v a l u e s would hold t o a h i g h e r a n g l e of a t t a c k as t h e s t a l l l e v e l i n c r e a s e d . Because some of t h e s p o i l e r p a n e l s were d e a c t i v a t e d f o r t h e PTA i n s t a l l a t i o n , i t was e s t i m a t e d t h a t s p o i l e r e f f e c t i v e n e s s would d i m i n i s h t o about 5 5 per- c e n t of t h e p u b l i s h e d G I 1 values. These d a t a i n d i c a t e , however, that PTA s p o i l e r e f f e c t i v e n e s s is about 62 p e r c e n t of t h e GI1 performance.

The e f f e c t of p r o p f a n power on t h e r o l l e f f e c t i v e n e s s of t h e s p o i l e r and a i l e r o n (one s i d e o n l y ) i s shown i n F i g u r e 5 9 . The a i l e r o n c o n t r i - b u t i o n was n o t a f f e c t e d by power, hence t h e d i f f e r e n c e s are e n t i r e l y s p o i l e r e f f e c t s . The l o c a l i n c r e a s e i n l i f t due t o the propfan s l i p s t r e a m c o n t r i b u t e d t o t h e i n c r e a s e d s p o i l e r e f f e c t i v e n e s s and t h e d e l a y i n t h e "drop-off" a t h i g h a n g l e of a t t a c k . The e f f e c t of s i d e s l i p on s p o i l e r r o l l power i s a l s o shown i n t h i s f i g u r e . A t n e g a t i v e s i d e s l i p a n g l e s ( r i g h t wing t r a i l i n g ) , t h e e f f e c t i v e n e s s of t h e s p o i l e r was reduced f o r both G I 1 and PTA c o n f i g u r a t i o n s and was l a r g e l y due t o t h e l o s s of wing l i f t on t h e t r a i l i n g wing. The propfan s l i p s t r e a m reduced t h e s i d e s l i p loss and hence improved the s p o i l e r e f f e c t i v e n e s s such t h a t e s s e n t i a l l y no change occurred w i t h s i d e s l i p . O p e r a t i o n a l l y , t h i s e f f e c t i s of l i t t l e consequence s i n c e , i n g e n e r a l , t h e s p o i l e r on t h e l e a d i n g wing would be d e f l e c t e d f o r s i d e s l i p c o n t r o l .

S p o i l e r e f f e c t i v e n e s s w i t h t a k e o f f f l a p is shown i n F i g u r e 60. The l i f t l o s s due t o s p o i l e r i n c r e a s e d because of t h e i n c r e a s e d wing l i f t from t h e f l a p s , and c o n s e q u e n t l y , t h e r o l l e f f e c t i v e n e s s of t h e s p o i l e r i n c r e a s e d . The f u l l - s p a n G I 1 s p o i l e r produced a r o l l i n g moment c o e f f i - c i e n t of 0.0344 compared t o t h e PTA level ( p r o p s o f f ) of 0.0194. Thus, t h e PTA s p o i l e r e f f e c t i v e n e s s was 56 p e r c e n t of t h e G I 1 l e v e l .

S i m i l a r d a t a f o r t h e l a n d i n g f l a p c o n f i g u r a t i o n a r e shown i n F i g u r e 61. S p o i l e r r o l l i n g moment e f f e c t i v e n e s s was f u r t h e r i n c r e a s e d r e l a t i v e t o t h e flaps-up l e v e l and i s 60 p e r c e n t of t h e published G I 1 l e v e l w i t h l a n d i n g f l a p .

The complete r o l l c o n t r o l c o n s i s t s of l e f t and r i g h t a i l e r o n d e f l e c - t i o n of f10 d e g r e e s and s p o i l e r d e f l e c t i o n on " a i l e r o n up" wing. The a i l e r o n e f f e c t i v e n e s s was unchanged by t h e PTA c o n f i g u r a t i o n and d o e s n o t vary w i t h f l a p d e f l e c t i o n . For low-speed f l i g h t , t h e t o t a l c o n t r i b u t i o n from t h e a i l e r o n t o t h e r o l l i n g moment c o e f f i c i e n t was 0.020. The r e s u l t a n t s p o i l e r / a i l e r o n e f f e c t i v e n e s s of t h e PTA c o n f i g u r a t i o n r e l a t i v e t o the published G I 1 l e v e l becomes 8 6 p e r c e n t , 76 p e r c e n t , and 7 1 p e r c e n t f o r t h e c l e a n , t a k e o f f , and l a n d i n g f l a p c o n f i g u r a t i o n s , r e s p e c t i v e l y .

O p e r a t i o n a l l y , t h e s e v a l u e s have t o be a d j u s t e d f o r f l e x i b i l i t y e f f e c t s , and s i n c e the PTA wings are c o n s i d e r a b l y s t i f f e r t h a n t h e GI1 wings, t h e p e r c e n t a g e l o s s of e f f e c t i v e n e s s due t o t h e reduced span s p o i l e r should b e lowered.

7.3.5 E f f e c t of Nacelle I n c i d e n c e - Low Speed

The e f f e c t of n a c e l l e i n c i d e n c e changes on l i f t and p i t c h i n g moment i s shown i n F i g u r e s 6 2 and 63. V i r t u a l l y no change i n l i f t occurred over t h e o p e r a t i o n a l range of a n g l e of a t t a c k . Small changes i n maxLmum l i f t C o e f f i c i e n t are apparent. These are probably due t o d i f f e r e n c e s i n wing l e a d i n g edge p r e s s u r e changes i n the v i c i n i t y of t h e n a c e l l e t h a t would probably n o t e x i s t a t f u l l - s c a l e Reynolds number. A s m a l l p o s i t i v e change CONFIGURATION R T AlLN R T SPLR D GI1 -10 -35 0 PTA (PROP OFF) -10 -35 A PTA ( T C = 0.1) -10 -35 + P T A ( T C = 0.9) -10 -35 FLAPS UP BETA=O FLAPS UP ALPHA=3 DEC -4 . 1 ( u u G- G- Z z t i !

w u -

v

U U.

U U w w 0 0 u u t- t - z Z 0 0 I I V f f J -I J J z z Rolling Moment From Aileron and Spoiler Deflections- Figure 5 9 .

Flaps Up

*-

CONFIGURATION R T SPLR 0 GI1 -35 O PTA (PROP (OFF) -35 FLAPS 20 A PTA (FEATHERED) -35 BETA = 0 4. I a.

-4. I I I I I I A T T A C K , CYlDEG) ANGLE OF

I 1

I

Figure 60.

Spoiler Effectiveness in Pitch - Flaps

20-Deg rees CONFIGURATION R T SPLR 0 G11 -35 0 PTA (PROP OFF) -35 FLAPS 1 1 0 A PTA (FEATHERED) -35 BETA = 0

4 v i

--la ANGLE OF ATTACK, ff (DEG) I I

I

z U ANGLE OF ATTACK (Y (DEG)

I I I I I I

I

I- v - z q SI-- s = w u v I i i -ILL - I W 0 0 a v 40-Deg rees

Figure 6 1 . Spoiler Effectiveness in Pitch - Flaps

NAC INDICENCE -1 DEG (BASELINE 0 -3 DEG D +2 DEG Effects of Nacelle Incidence on Lift and Pitching Figure 6 2 .

Moment - Flaps Up J TC NAC INCIDENCE 1.8 0 . 1 - 1 DEG (BASELINE) 1 7 I, $1 - 3 DEG FLAPS UP ,I *I +2 DEG n + 0.85 0.9 -1 DEG (BASELINE) I, 1' -3 DEG X ,I I, +2 DEG Effects of Nacelle Incidence on Lift and Pitching Figure 6 3 .

Moment - Flaps Up, Propfan Power On

i n p i t c h i n g moment o c c u r s w i t h +2 d e g r e e s of i n c i d e n c e r e l a t i v e t o -1 degree t h a t i s a f u n c t i o n of a n a c e l l e l i f t change. There Ls a l s o a n e g a t i v e p i t c h i n g moment change w i t h -3 d e g r e e s t h a t has been masked by a similar p o s i t i v e change a t t h e h o r i z o n t a l t a i l . I n t h e presence of prop- f a n power, t h e s e p i t c h i n g moment changes are s t r o n g e r and more uniform.

The t a i l changes a r e small and random and i n s i g n i f i c a n t . The t r i m changes a s s o c i a t e d w i t h n a c e l l e i n c i d e n c e are small b u t uniform.

The e f f e c t of t h e n a c e l l e i n c i d e n c e changes on l a t e r a l - d i r e c t i o n a l d a t a i n s i d e s l i p i s v e r y small and of no consequence, The change i n side- f o r c e w i t h n a c e l l e i n c i d e n c e i s t h e o n l y measurable parameter w i t h h i g h propfan power and should c a u s e small changes i n bank a n g l e f o r t r i m t h a t would be d i f f i c u l t t o d e t e c t .

7.3.6 Lift and Pitch - High Speed

The v a r i a t i o n of l i f t and p i t c h i n g moment w i t h Mach number f o r t h e G I 1 and v a r i o u s PTA c o n f i g u r a t i o n s i s shown i n F i g u r e s 64 through 67 f o r Mach numbers from 0.4 t o 0.85. A l i f t increment due t o t h e PTA n a c e l l e , of t h e same magnitude as t h e low-speed d a t a , can be seen i n Figure 64 a t lower a n g l e s of attack. A s a n g l e of a t t a c k i n c r e a s e d , t h i s l i f t increment changed t o a l o s s as a r e s u l t of i n n e r wing l e a d i n g edge l o a d i n g s h i f t s .

This l o a d s h i f t a l s o produced a s t a b l e change i n p i t c h i n g moment a t high a n g l e s of a t t a c k t h a t are above t h e a n g l e s f o r o p e r a t i o n a l needs. For l i f t c o e f f i c i e n t s between 0.2 and 0.4, t h e p i t c h i n g moment d i f f e r e n c e s between t h e PTA and G I 1 c o n f i g u r a t i o n s became smaller as Mach number is i n c r e a s e d t o 0 . 8 5 . Hence only s m a l l d i f f e r e n c e s i n t r i m requirements and l i f t a t c o n s t a n t a n g l e of a t t a c k are t o be expected.

Composite p i c t u r e s showing c o m p r e s s i b i l i t y e f f e c t s on l i f t d a t a f o r t h e G I 1 and PTA c o n f i g u r a t i o n s are shown i n F i g u r e 68. These p l o t s were c o n s t r u c t e d u s i n g t h e low-speed and high-speed test r e s u l t s . For t h e s e p l o t s , t h e low-speed d a t a , which were o b t a i n e d a t a Reynolds number of 1.65 m i l l i o n , have been c o r r e c t e d t o t h e Reynolds number of t h e high-speed d a t a (3.35 m i l l i o n based on wing chord). The l i f t c o r r e c t i o n i s p r i m a r i l y due t o wing l i f t - c u r v e - s l o p e changes.

It can be s e e n i n F i g u r e 68 t h a t t h e PTA maximum l i f t was s l i g h t l y reduced, and t h e c o m p r e s s i b i l i t y e f f e c t peaked a t a s l i g h t l y lower Mach number than f o r t h e G I I . It i s t h e s e two parameters t h a t d e f i n e t h e b u f f e t o n s e t curve f o r t h e o p e r a t i o n a l speed range. Superimposed on t h e G I 1 l i f t d a t a i s t h e b u f f e t o n s e t curve d e r i v e d from f l i g h t experience.

A t low Mach numbers, t h i s curve i s p r i m a r i l y a n i n d i c a t i o n of t h e begin- ning of s t a l l s e p a r a t i o n ; a t h i g h Mach numbers, it i s a r e s u l t of shock s h i f t s and boundary l a y e r t h i c k n e s s adjustments c a u s i n g wing load changes t h a t r e s u l t i n b u f f e t i n g . T h i s l a t t e r u s u a l l y becomes n o t i c e a b l e c l o s e t o t h e l i f t peak a t a p a r t i c u l a r Mach number and a n g l e of a t t a c k combination as i n d i c a t e d by t h e 2- and 3-degree points. A t 4 and 5 d e g r e e s the test d a t a i n d i c a t e a b u f f e t o n s e t p r i o r t o t h e peak l i f t f o r t h e G I I . The PTA c o m p r e s s i b i l i t y peak occurred a l i t t l e earlier and would t h e r e f o r e i n d i c a t e a n earlier b u f f e t onset.

and p i t c h i n g moment d a t a from low- 70 summarize l i f t F i g u r e s 69 and and high-speed tests. The e f f e c t s of c o m p r e s s i b i l i t y on t h e p i t c h i n g moment a t c o n s t a n t l i f t , as shown i n F i g u r e 69, i n d i c a t e a t y p i c a l Mach CONFIGURATION 0 GI1 MACH = 0 . 4 0 0 GI1 + P I A NACELLE A PTA LESS ACOUSTIC BOOM + PTA

PTA Configuration Buildup - Mach 0 . 4

Figure 64.

c

*-

CONFIGURATION 0 G I I O G I I + PTA NACELLE Mach 0.70 A P T A LESS ACOUSTIC BOOM + PTA Figure 65.

PTA Configuration Buildup - Mach 0.7

CONFIGURATION 0 G I I 0 GI1 + PTA NACELLE = 0.80 PTA LESS ACOUSTIC BOOM MACH 4 - PTA

PTA Configuration Buildup - Mach 0 . 8

Figure 66.

*-

CONFIGURATION 0 GI1 OGll + PTA NACELLE MACH = 0.85 A P T A LESS AGOUSTlC BOOM 4- PTA ~ ~ - 0 . 2 I ANGLE OF A T T A C K , a (DEC)

Figure 6 7 . PTA Configuration Buildup - Mach 0 . 8 5

C -

*-

GI I-CONFIGURATION 1 . 2 1 . 0 cL 0 . 8 0 . 6 0.4 0 . 2 P T A CONFIGURATION Mach Number Effects on Lift Characteristics Figure 68.

F K I 0 . 4 c f :: A I I I I I I a I I I- W z 0 0.4 0 . 6 0 . 8 MACH NO.

0 G I I E l PTA PROP. OFF X P T A - T z U I I I I I I I I I- z W z V f r V t a +

u--

A MACH NO.

n: W N -0.04 ~ Figure 69. Comparison of Lift Data From Low- and High-speed Test 0 GI1 PTA PROP. OFF Tc MAX 0.2 E J u I 0.1 0 0.4 0.6 0 . 2 0 . 8 MACH NO.

- 2 . 0 -1.0 u .

0.8 0 0 . 2 0.4 0.6 MACH NO.

Test Figure 70. Comparison of Pitch Data From Low- and High-speed "tuck" above a Mach number of 0.8 f o r both t h e G I 1 and t h e PTA configura- t i o n s . This nose down tendency is a l i t t l e more pronounced f o r t h e PTA c o n f i g u r a t i o n . The Mach t r i m compensator i n t h e G I 1 c o n t r o l system should be adequate f o r t h e PTA as w e l l .

The l i f t change due t o t h e PTA i n s t a l l a t i o n ( F i g u r e 70) results i n a change i n angle of a t t a c k f o r z e r o l i f t w i t h o n l y minor changes i n l i f t curve slope. The l i f t i n c r e a s e due t o power a t low speeds r e s u l t s i n changes i n both. The d e c r e a s e i n a n g l e of a t t a c k f o r z e r o l i f t r e s u l t s i n an i n c r e a s e i n p i t c h i n g moment a t z e r o l i f t c o n t r i b u t e d by t h e t a i l as can be s e e n i n Figure 69 a t M = 0.4. A t t h e low-speed t e s t p o i n t , t h e apparent lower t a i l i n p u t t o p i t c h i n g moment a t z e r o l i f t i s due t o t h e lower Reynolds number a t t h e t a i l as p r e v i o u s l y discussed. The forward movement of t h e n e u t r a l p o i n t due t o t h e PTA c o n f i g u r a t i o n is s e e n t o be uniform over t h e Mach range. The e f f e c t of power i s most pronounced a t low speeds (high Tc).

7.3.7 L a t e r a l - D i r e c t i o n a l E f f e c t s - High Speed

L a t e r a l - d i r e c t i o n a l d a t a are shown i n F i g u r e s 71 through 74 i n terms of s i d e f o r c e , yawing moment, and r o l l i n g moment c o e f f i c i e n t s at *5 degrees of s i d e s l i p as a f u n c t i o n of angle of a t t a c k . No d i r e c t s i d e s l i p d a t a a t c o n s t a n t a n g l e of a t t a c k were o b t a i n e d i n t h e high-speed test due t o t u n n e l support c o n s t r a i n t s . The o f f s e t s a t z e r o s i d e s l i p are a l s o shown. The s i d e f o r c e and r o l l i n g moment e f f e c t s f o r t h e GI1 are a r e s u l t of t h e aerodynamic l o a d s on t h e non-metric cover a t t h e model-to-sting i n t e r f a c e . While i t i s impossible t o f u l l y e x t r a c t t h i s increment from t h e d a t a , s i n c e t h e cover l o a d s vary as a f u n c t i o n of t h e c o n f i g u r a t i o n , i t i s f e l t t h a t t h e i n c r e m e n t a l e f f e c t of t h e PTA w i l l be r e p r e s e n t a t i v e .

From Figure 7 1 , a t Mach 0.4, t h e i n c r e m e n t a l e f f e c t of s i d e s l i p on s i d e f o r c e i s seen t o i n c r e a s e , r e l a t i v e t o t h e G I I , when t h e n a c e l l e is l e a d i n g and d e c r e a s e , r e l a t i v e l y , when t h e n a c e l l e i s t r a i l i n g . Yawing moment due t o s i d e s l i p shows l i t t l e v a r i a t i o n w i t h a n g l e of a t t a c k and only a s l i g h t r e d u c t i o n r e l a t i v e t o t h e G I I . Roll due t o s i d e s l i p i s s l i g h t l y reduced by t h e PTA c o n f i g u r a t i o n b u t has t h e same rate of i n c r e a s e with angle of a t t a c k as t h e GII.

S i m i l a r e f f e c t s are s e e n i n Figure 72 a t a Mach number of 0.7 and Figure 73 a t a Mach number of 0.8. A t a Mach number of 0.85, F i g u r e 74, t h e r e d u c t i o n i n roll due t o s i d e s l i p a t h i g h a n g l e s of a t t a c k i s s t r o n g e r f o r t h e PTA than f o r t h e G I 1 and may be n o t i c e a b l e t o t h e p i l o t i f he a t t e m p t s t o p i c k up a wing w i t h c o n t r o l l e d s i d e s l i p . It may a l s o be noted t h a t t h e o f f s e t i n r o l l a t z e r o s i d e s l i p i s reduced t o z e r o a t small a n g l e s of a t t a c k and becomes p o s i t i v e a t h i g h e r angles. This has t h e e f f e c t of making t h e o f f s e t i n weight due t o t h e n a c e l l e appear h e a v i e r a t h i g h Mach numbers.

A summary of t h e s i d e s l i p d e r i v a t i v e s and t h e o f f s e t s a t z e r o s i d e - s l i p are shown as f u n c t i o n s of Mach number i n F i g u r e s 75 and 76. It can be s e e n t h a t t h e r e i s very l i t t l e d i f f e r e n c e i n t h e d e r i v a t i v e s r e l a t i v e t o t h e G I 1 of s i g n i f i c a n c e t o a i r c r a f t h a n d l i n g q u a l i t i e s for t e s t b e d o p e r a t i o n s o t h e r than t h e r e d u c t i o n i n roll due t o s i d e s l i p . T h i s reduc- t i o n t e n d s t o i n c r e a s e t h e d u t c h roll s t a b i l i t y and i n c r e a s e t h e s p i r a l i n s t a b i l i t y . This e f f e c t i s of l i t t l e consequence e s p e c i a l l y w i t h a working yaw damper.

CONFIGURATION BETA 0 GI1 +5 DEG 0 " 0 DEG MACH = 0 . 4 0 0 n " - 5 DEG 4- PTA +5 DEC PROP OFF 0 DEG - 5 DEG C V i V c z c z V z w w

u

G-

u

U Z U U U.

w W W

u

V U V U W k-

- 5

Z W V W ANGLE OF ATTACK, z I W V I I a : V V f z -I W I -I D d 0 v) > a :

Figure 7 1 . Comparison of GI I and PTA in Sideslip - Mach 0 . 4

CONFIGURATION BETA 0 CII +5 DEC MACH = 0 . 7 0 0 (1 0 DEG a " - 5 DEG PROP OFF + PTA +5 DEG x 43 0 DEC 0 " -5 DEG ANGLE OF ATTACK a (DEG 4NGLE OF A T T A C K , a (DEG)

Figure 7 2 . Comparison of GI1 and PTA in Sideslip - Mach 0 . 7

*-

CONFIGURATION BETA 0 GI1 +5 DEG 0 DEG 0 " - 5 DEG a " MACH = 0.80 f PTA +5 DEG PROP OFF x $1 0 DEG - 5 DEG 0 "

Figure 73. Comparison of CII and PTA in Sideslip - Mach 0 . 8

*

*-

CONFIGURATION BETA 0 GII +5 DEG 0 DEC 0 " a " -5 DEC MACH = 0 . 8 5 4- PTA +5 DEC PROP OFF x $1 0 DEC 0 " -5 DEC V7 C" I-'

< z

z w w I!

U I!

U .

U U W W 0 0 V V I- I - z W I I V z

-

-I -I lY

Figure 7 4 . Comparison of GI1 and P T A in Sideslip - Mach 0 . 8 5

(I = 30

+ NACELLE LEADING

X NACELLE T R A I L I N G 0 GI1 - BASE -0.02 C I I I I I I I I I I 0 0 . 2 0 . 4 0.6 0.8 1 . o MACH NUMBER 0.004 a C “ P 0 - 0.2 0 . 4 0.6 0.8 1.0 MACH NUMBER -0.002 4, K I + 0 0.2 0.4 0.6 0.8 1.0 MACH NUMBER Figure 75. Effect of Mach Number on Sideslip Derivatives 1 0 1 n = 30 PROP OFF 0.1 U 0.01

-0.01 I

0 . 0 1 (Acl’ p = 0 0 -0.01 0 0.2 0.4 0 . 6 0.8 1.0 MACH NUMBER Figure 76. Effect o f Mach Number on Lateral-Directional Offsets

7.3.8 E l e v a t o r E f f e c t i v e n e s s - High Speed

No e l e v a t o r d a t a were obtained i n t h e high-speed wind t u n n e l tests s i n c e no change i n e f f e c t i v e n e s s w a s p r e d i c t e d . A p o s s i b l e change i n t a i l a n g l e of a t t a c k due t o t h e n a c e l l e , or l o c a l dynamic p r e s s u r e changes, would have been t h e o n l y s o u r c e of a change i n e l e v a t o r e f f e c t i v e n e s s .

The t a i l a n g l e of a t t a c k change due t o t h e n a c e l l e i s less t h a n 0.5 d e g r e e s , and no l o c a l dynamic p r e s s u r e change was d e t e c t e d .

7.3.9 Rudder E f f e c t i v e n e s s - High Speed

The e f f e c t i v e n e s s of t h e rudder a t small d e f l e c t i o n s ( 5 d e g r e e s ) i s shown i n Figure 77 as a f u n c t i o n of Mach number. These shapes are t y p i c a l f o r a swept f i n a i r c r a f t . The d a t a show a h i g h e r l e v e l of e f f e c t i v e n e s s f o r t h e G I 1 t h a n i s published t o g e t h e r w i t h a w e l l d e f i n e d c o m p r e s s i b i l i t y e f f e c t . This is t h e only a r e a of s i g n i f i c a n t disagreement w i t h published s t a b i l i t y and c o n t r o l d a t a i n t h e whole test series. It i s p o s s i b l e t h a t t h e h i g h e r e f f e c t i v e n e s s l e v e l i s due t o non-metric l o a d s on t h e s t i n g cover. However, d a t a obtained from t h e QUADPAN model a t M = 0.45 and M = 0.8 tend t o v e r i f y t h e l e v e l as measured. The PTA d a t a , power on, shows a s i g n i f i c a n t i n c r e a s e i n rudder e f f e c t i v e n e s s r e l a t i v e t o t h e G I 1 a t M = 0 . 4 and 0.7. A t M = 0.8, an e a r l i e r break i n t h e c o m p r e s s i b i l i t y r i s e r e s u l t s i n a s l i g h t r e d u c t i o n i n rudder e f f e c t i v e n e s s . These changes would have l i t t l e e f f e c t on t h e a i r c r a f t o p e r a t i o n .

7.3.10 Aileron-Spoiler E f f e c t i v e n e s s - High Speed

A i l e r o n and s p o i l e r e f f e c t i v e n e s s from M = 0 . 4 t o M = 0 . 8 5 is shown i n Figure 78. The a i l e r o n d e f l e c t i o n of 10 d e g r e e s , t r a i l i n g edge up, i s t h e maximum t r a v e l . A t low Mach numbers, t h e r o l l i n g moment due t o 10 d e g r e e s of a i l e r o n i s 7-percent h i g h e r than t h e G I 1 published l e v e l . A t h i g h Mach number an e a r l y break i n c o m p r e s s i b i l i t y e f f e c t occurs and r e s u l t s i n lower e f f e c t i v e n e s s beyond a Mach number of 0.7. The wind t u n n e l G I 1 d a t a has a s i m i l a r Mach number e f f e c t as shown f o r an a n g l e of a t t a c k of 3 degrees. The f u l l - s c a l e a i r p l a n e h a s v o r t e x g e n e r a t o r s on t h e upper s u r f a c e of t h e wing outboard of t h e f e n c e t o t a k e care of t h i s phenomenon, and i t i s t h e r e f o r e assumed t h e PTA f u l l - s c a l e w i l l have a t l e a s t t h e GI1 f u l l - s c a l e e f f e c t i v e n e s s . S i m i l a r l y , s p o i l e r e f f e c t i v e n e s s i s seen t o be h i g h e r than t h e 55 p e r c e n t of t h e f u l l - s c a l e G I 1 l e v e l esti- mated f o r t h e PTA s h o r t span s p o i l e r b u t w i t h a n e a r l y shock r e l a t e d loss.

The v o r t e x g e n e r a t o r s on t h e f u l l - s c a l e a i r c r a f t are expected t o r e s t o r e t h i s l e v e l . Maximum s p o i l e r d e f l e c t i o n is nominally 35 d e g r e e s as shown i n t h e low-speed s e c t i o n and is a v a i l a b l e a t h i g h speed.

7 . 4 LEX (LEADING EDGE EXTENSION) PERFORMANCE The o b j e c t i v e s of t h e LEX d e s i g n and t h e methodology t o a c h i e v e those o b j e c t i v e s are d e s c r i b e d i n Appendix A. G e n e r a l l y , i t was d e s i r e d t o recontour t h e wing l e a d i n g edge r e g i o n s o t h a t it would be less s e n s i t i v e t o t h e upwash produced by t h e propfan s l i p s t r e a m .

Wind tunnel test d a t a showing t h e e f f e c t s of t h e LEX on wing p r e s s u r e d i s t r i b u t i o n s are shown i n F i g u r e s 79 and 80. Data are shown a t Mach G11 PUBLISHED e GI1 WIND TUNNEL

-+- PTA WIND TUNNEL

MAX Tc x QUADPAN GI1 t PTA 0.4 0.6 0.8 0 . 4 0 . 6 0.8 0.4 0.6 0 . 8 MACH NO.

MACH NO. MACH NO.

Figure 77. Comparison of Measured and Predicted Rudder Effectiveness

SPOILER - 10"

" V V " I I L 0.4 0 . 6 0 . 8 1 .o MACH NO.

Figure 78. Aileron-Spoiler Effectiveness 1 7 = .290. n = 2 O . M = 0.7 1 0 * LOWER SURFACE I S SHADED Figure 79. Effect o f LEX on Wing Surface Pressures, Mach 0.70 -1.60 0 PROP OFF

a PTA

4- PTA LEX -1.20

t

-0.80 A -0- -0.40 /+- CP

l-

-0.00

0.1- 0

1.00 X I C 0.40 0.80 1.20 * LOWER SURFACE I S SHADED Effect of LEX on Wing Surface Pressures, Mach 0.80 Figure 80.

numbers of 0.7 and 0.8 f o r t h e PTA b a s i c c o n f i g u r a t i o n w i t h and without t h e propfan and f o r t h e powered c o n f i g u r a t i o n w i t h t h e LEX added. Unfor- t u n a t e l y , t h e r e were no p r o v i s i o n s i n t h e LEX hardware i t s e l f f o r p r e s s u r e measurements, so t h e first few p o i n t s of t h e p r e s s u r e d i s t r i b u t i o n curves are missing f o r t h e LEX c o n f i g u r a t i o n . There are some c o n c l u s i o n s t h a t can be drawn, however, w i t h t h e p o i n t s a v a i l a b l e .

The p r e s s u r e d i s t r i b u t i o n s shown i n Figure 79 f o r t h e c o n f i g u r a t i o n w i t h t h e LEX have t h e chord dimension a d j u s t e d f o r t h e a d d i t i o n a l l e n g t h of t h e LEX. A t Mach 0.7, i t can be seen t h a t t h e e f f e c t of power on t h e b a s i c GI1 wing i s t h e same as an angle-of-attack increase--there i s a s l i g h t i n c r e a s e i n Mach number (more n e g a t i v e Cp) and t h e shock wave is s l i g h t l y f u r t h e r a f t . For t h e same power c o n d i t i o n w i t h t h e LEX, however, t h e r e i s c o n s i d e r a b l y less a c c e l e r a t i o n over t h e f r o n t p a r t of t h e air- f o i l . I n t h i s case, t h e LEX h a s c l e a r l y done what it w a s designed t o do.

A t Mach C . 8 , t h e a c c e l e r a t i o n over t h e l e a d i n g edge is not a s s t r o n g w i t h t h e LEX as without i t , b u t t h e flow does accelerate t o a h i g h e r Mach number. While t h i s could have had a n e g a t i v e e f f e c t , t h i s does not appear t o be t h e case because t h e d e c e l e r a t i o n behind t h e peak i s mild w i t h no evidence of t r a i l i n g edge s e p a r a t i o n as t h e r e was i n t h e case without t h e LEX.

A d r a g p o l a r p l o t f o r t h e Mach 0.8 case is shown i n F i g u r e 81. This p l o t confirms t h e s u c c e s s of t h e LEX design--showing t h a t t h e a d d i t i o n of t h e LEX reduced d r a g a t t h e design l i f t c o e f f i c i e n t by approximately 29 counts. A t h i g h e r l i f t c o e f f i c i e n t s , t h e e f f e c t of t h e s o f t e n e d l e a d i n g edge of t h e LEX c o n f i g u r a t i o n i s even g r e a t e r ; whereas a t lower l i f t coef- f i c i e n t s , t h e b a s i c c o n f i g u r a t i o n does n o t g e t i n t o as much t r o u b l e , and t h e b e n e f i t s from t h e LEX are less.

7.5 WAKE SURVEY DATA 7.5.1 Wing P r e s s u r e Measurements It was d e s i r e d i n t h e flow survey tests t h a t flow about t h e semispan model match as c l o s e l y a s p o s s i b l e t h e flow about t h e f u l l - s p a n PTA model.

To check t h a t t h i s w a s accomplished, comparisons were made of t h e wing p r e s s u r e d i s t r i b u t i o n s from t h e flow f i e l d tests w i t h t h o s e o b t a i n e d i n t h e 16-Ft Transonic Aerodynamics Wind Tunnel f o r t h e f u l l - s p a n model.

These comparisons are shown i n F i g u r e s 82 and 8 3 f o r two wing spanwise s t a t i o n s .

I n Figure 8 2 , f o r q = , 3 2 8 , t h e d a t a from t h e flow f i e l d survey tests a t z e r o a n g l e of a t t a c k a r e compared w i t h d a t a from t h e 16-Ft Transonic Aerodynamics Wind Tunnel a t z e r o and 1-degree a n g l e s of a t t a c k . It can be s e e n t h a t t h e flow survey t e s t d a t a on t h e upper s u r f a c e f a l l g e n e r a l l y between t h e two d a t a p l o t s from t h e o t h e r tunnel. T h i s i m p l i e s t h a t t h e r e i s an e f f e c t i v e angle-of-attack d i f f e r e n c e between t h e two f a c i l i t i e s of a degree or less.

t h e o r d e r of h a l f On t h e lower s u r f a c e , t h e r e w a s a m o d i f i c a t i o n of t h e flow survey model t o accommodate r a k e i n s t r u m e n t a t i o n t u b e s e x i t i n g t h e model. This produced a protuberance t h a t caused t h e p r e s s u r e d i s t r i b u t i o n s t o be d i f - f e r e n t from those i n t h e f u l l span model tests. However, i n t h e l e a d i n g edge r e g i o n , t h e flow survey test d a t a s t i l l f a l l between t h e o t h e r two c u r v e s .

*-

MACH 0.8 0 LEX O N 0 LEX OFF 0.5 0.4 cL 0 . 3 0 . 2 .02 .03 .04 .05 .06 cD Figure 8 1 . Effect o f LEX on PTA Drag -1.60 f M = 0.70 rl = 0.328 a TUNNEL DOo LANGLEY 16 F T 00' LEWIS 8 F T Y 6 F T -1.20 A l u LANGLEY 16 FT fl DENOTES LOWER SURFACE -0.80 C P -0.40 -0.00 0.80 1.20 Figure 82. Pressure Distributions Inboard of Nacelle M = 0.70 r) = 0.511 a TUNNEL 0 Oo LANGLEY 16 FT -1.20 0 O o LEWIS 8 FT X 6 FT A l o LANGLEY 16 FT a DENOTES LOWER SURFACE - 0 . EO -0.40 C P -0.00 AI - 0.40

0.80 -

1.20 -

Figure 83. Pressure Distributions Outboard of Nacelle On t h e outboard s i d e of t h e n a c e l l e , t h e d a t a i n F i g u r e 83 e x h i b i t t h e same c h a r a c t e r i s t i c s a s t h a t s e e n i n F i g u r e 82. It w a s concluded then t h a t the flow f i e l d s around t h e semispan model were b a s i c a l l y t h e same as t h o s e around t h e f u l l span model, b u t that t h e a n g l e of a t t a c k i n t h e flow f i e l d tests i s e f f e c t i v e l y about one-half degree h i g h e r than t h a t i n t h e f u l l - s p a n model tests.

7 . 3 . 2 Flow Field Data The use of E h o l e probes t o measure three-dimensional v e l o c i t y v e c t o r s is s t i l l somewhat a n a r t - - p a r t i c u l a r l y i n high speed flows. Con- s e q u e n t l y , some f a c t o r s i n t h e s e tests r e s u l t e d i n unexpected e f f e c t s on t h e data. Looking a t F i g u r e 6 , two f e a t u r e s of t h e experimental set-up should be noted. F i r s t , t h e rakes were i n s t a l l e d as c l o s e as p o s s i b l e t o t h e i n l e t f a c e on t h e model. This was done i n i t i a l l y when it was thought t h a t t h e s e tests would i n c l u d e powered e f f e c t s ; t h u s , t h e rakes were mounted s o t h a t one p o s i t i o n would s u f f i c e f o r both prop-on and prop-off c o n f i g u r a t i o n s .

The second f e a t u r e t o be noted i s t h e attachment of t h e rake t o t h e model. The bracket t o which t h e rake a t t a c h e d presented a flow i n t e r f e r - ence g r e a t e r t h a n d e s i r a b l e i n a t r a n s o n i c flow. The e f f e c t s of both t h e i n l e t l i p s and t h e support b r a c k e t were measured by t h e rake probes (as t h e y should be). The p r e d i c t i o n code, QUADPAN, on t h e o t h e r hand, w a s not paneled i n s u f f i c i e n t d e t a i l t o a d e q u a t e l y model t h e s e f e a t u r e s of t h e model hardware. Thus, as w i l l be seen l a t e r , t h e p r e d i c t i o n s f a i l e d t o r e f l e c t a l l of t h e flow nuances d e t e c t e d by t h e rakes.

Flow f i e l d d a t a f o r Mach 0 . 6 are presented i n F i g u r e s 84 through 86 i n t h e form of v a l u e s of t h e t h r e e v e l o c i t y components U , V , and W normalized w i t h respect t o f r e e s t r e a m v e l o c i t y . In each f i g u r e d a t a are p r e s e n t e d f o r t h e r a k e s a t f o u r azimuthal p o s i t i o n s .

I n Figure 84, v a l u e s are shown f o r t h e axial component U. The agreement of t h e d a t a w i t h t h e Q U A D P A N p r e d i c t i o n s is not very good, I n F i g u r e s 85 and 86, t h e d a t a f o r t h e v e r t i c a l component W and t h e l a t e r a l component V show much b e t t e r agreement w i t h QUADPAN p r e d i c t e d values. It i s n a t u r a l t o q u e s t i o n whether t h e V and W d a t a are c r e d i b l e when t h e U d a t a show such poor agreement. It is b e l i e v e d , however, t h a t t h e V and W d a t a should be accepted f o r t h e following reasons. F i r s t , t h e measurement of U i s an a b s o l u t e value; t h e measurements of V and W, on t h e o t h e r hand, are r e l a t i v e - t h a t is, t h e y depend on d i f f e r e n c e s between p r e s s u r e s measured a t d i f f e r e n t p o i n t s on t h e 5-hole probe. These r e l a t i v e measure- ments are i n h e r e n t l y more a c c u r a t e than t h e a b s o l u t e . Therefore, even i f t h e U v a l u e i s i n e r r o r , V and W may be c o r r e c t .

The second reason is r e l a t e d . It has a l r e a d y been p o i n t e d o u t t h a t t h e rakes are c l o s e r than d e s i r e d t o support and o t h e r model hardware.

This proximity not only c a u s e s t h e rake t o be i n r e g i o n s of l o c a l acceler- a t i o n and d e c e l e r a t i o n t h a t are not p r e d i c t e d by QUADPAN, b u t it a l s o may produce s p u r i o u s readings because t h e rake support i s not e x a c t l y t h e same a s t h a t used i n t h e c a l i b r a t i o n .

For t h e s e reasons, t h e approach taken i n t h e d i s c u s s i o n of t h e wake survey r e s u l t s has been t o d i s c o u n t , somewhat, t h e d a t a f o r t h e a x i a l component U and t o base c o n c l u s i o n s p r i m a r i l y on t h e d a t a f o r V and W.

MACH 0.6 N A C E U E ALPHA - 1.

ANGLE EXPERIMENT OF OUADPAN SYMBOL ATTACK SYMBOL

e -1 -

8 0 * - I - -

0 2 ------

0 I ---

A 6 -----

= On 1.2 1.1 > c g 1.0 i Y > a 0.9 0.8 ; I 0.0 0.2 0.u 0.6 0.8 1.0 1.1 NON-OIMENSIONAL PROP RADIUS NON-DIMENSIONAL PROP RADIUS _I = 1110.

I 270' 1 . 1 > c 3 1.0 Y a 0.9 0.8 0.0 0.2 0.6 0.6 0.1 1.0 1.2 NON-DIMENSIONAL PROP RADIUS NON-DIMENSIONAL PROP RADIUS Rake Data for Axial Velocity Component U at Mach 0 . 6 Figure 84.

MACH 0.6 NACELLE ALPHA - 1'

ANGLE EXPERIMENT OF QUADPAN SYMBOL ATTACK s Y M a o L

.-----. -2 -

n 0 --

------

- 4 --

-----

0 - 6 NOM-OIYENSIONAL PROP RADIUS NOH-OIMENSIONAL PROP RADIUS NACELLE ALPWA - -1.

Y l C Y 0 ' 0 . 1 0.1 e 0.0 U > 3 -0.1 -0.1 -0.1 WON-OIMENSIONAL PROP RADIUS NON-DIMENSIONAL PROP RADIUS Rake Data for Vertical Velocity Component W at Mach 0.6 Figure 8 5 .

*-

MACH 0.6 NACELLE ALPHA - I.

ANGLE EXPERIMENT OF OUAOPAN SYMBOL ATTACK SYMBOL

-1 -

0 -

0 a --____

0 8 --

A 6 -----

> .

0 . 3 > c 5 0.2 Y > > 0.1 0.0 -0.1 0.0 0.2 0.P 0.6 0.8 1.0 1.1 0.0 0.2 0.9 0.6 0.8 1.0 1.2 NON-DIMENSIONAL PROP RADIUS NON-DIMENSIONAL PROP RADIUS Figure 86. Rake Data for Lateral Velocity Component V at Mach 0.6 Data f o r t h e l a t e r a l component V show a much b e t t e r s e l f - c o n s i s t e n c y than those f o r U and a much b e t t e r c o r r e l a t i o n w i t h t h e QUADPAN predic- t i o n s . I n f a c t , o v e r a l l , t h e agreement w i t h p r e d i c t e d v a l u e s is good except f o r t h e rake p o s i t i o n i n f r o n t of t h e i n l e t ( J = 90 degrees).

Here, t h e disagreement is a t t r i b u t e d t o t h e inadequacy of t h e Q U A D P A N pane l i n g .

W i n Figure 86 show an e x c e l l e n t The d a t a for t h e v e r t i c a l component agreement w i t h t h e p r e d i c t e d values. A t a l l f o u r rake p o s i t i o n s t h e v a r i a t i o n of W w i t h d i s t a n c e from t h e prop c e n t e r l i n e and t h e v a r i a t i o n w i t h a n g l e of a t t a c k are w e l l p r e d i c t e d by QUADPAN.

Data recorded a t Mach 0 . 8 5 are shown f o r t h e t h r e e v e l o c i t y compo- n e n t s i n Figures 87 and 88. The t r e n d s observed i n t h e lower Mach number d a t a are g e n e r a l l y repeated i n t h e s e f i g u r e s . Obviously, t h e r e is no breakdown i n t h e QUADPAN code a t t h e higher Mach number.

were performed f o r t h e purpose of These flow f i e l d measurement tests demonstrating t h a t t h e subsonic i n v i s c i d code QUADPAN can be used t o p r e d i c t t h e propfan flow f i e l d f o r a l l t h e c o n d i t i o n s of t h e PTA f l i g h t t e s t program. I n s p i t e of t h e d i f f i c u l t i e s experienced, i t i s believed t h a t t h e d a t a provide t h i s demonstration.

7 . 6 ISOLATED PROPELLER TEST The g o a l i n t h e i s o l a t e d p r o p e l l e r tests w a s t o assess t h e p r o p e l l e r performance i n t h e absence of i n s t a l l a t i o n e f f e c t s . I n o r d e r t o do t h i s , t h e p r o p e l l e r should, i d e a l l y , have a support n a c e l l e no l a r g e r than t h e hub diameter of t h e p r o p e l l e r . Because t h i s support must e n c l o s e t h e p r o p e l l e r d r i v e system and a l l t h e n e c e s s a r y i n s t r u m e n t a t i o n t o monitor p r o p e l l e r performance, t h i s i d e a l was i m p r a c t i c a l , so an axisymmetrical n a c e l l e a s small a s p o s s i b l e was employed.

T e s t s were run a t Mach 0.4 with t h e t h r u s t a x i s a n g l e of a t t a c k ranging from -2 t o 1 2 d e g r e e s and p r o p e l l e r b l a d e p i t c h set a t 49 degrees (measured a t t h e 3/4-blade r a d i a l s t a t i o n ) . S i x advance r a t i o s were s e t by varying t h e p r o p e l l e r r o t a t i o n a l speed from 11,000 t o 16,000 rpm.

V i b r a t i o n problems precluded t e s t i n g a t Mach numbers h i g h e r than 0.4 The d a t a obtained from t h e hub balance included p r o p e l l e r a x i a l f o r c e , normal f o r c e , and p i t c h i n g moment. Using t h e a i r motor d r i v e p r e s s u r e , p r o p e l l e r r o t a t i o n a l speed, and previous c a l i b r a t i o n d a t a , t h e torque and horsepower absorbed by t h e p r o p e l l e r were deduced. These d a t a allowed t h e d e t e r m i n a t i o n of p r o p e l l e r performance and its v a r i a t i o n w i t h p r o p e l l e r advance r a t i o and a n g l e of a t t a c k .

Thrust and power c o e f f i c i e n t s , p l o t t e d a g a i n s t p r o p e l l e r advance r a t i o r e f l e c t p r o p e l l e r performance. These a r e p l o t t e d i n F i g u r e s 89 and 90. The d a t a a r e i n good agreement w i t h a n a l y t i c a l p r e d i c t i o n and a l s o c o r r e l a t e w e l l w i t h t h a t p r e d i c t e d f o r t h e f u l l - s c a l e p r o p e l l e r .

Data from t h e i s o l a t e d p r o p e l l e r test i n t h e Langley 4 M x 7M Subsonic Wind Tunnel a r e shown i n F i g u r e s 91 and 92. Data are shown f o r blade p i t c h a n g l e s of 38 degrees and 40 degrees. C o r r e l a t i o n w i t h theory h e r e is n o t as good a s i n t h e high-speed tests. One reason f o r t h i s discrepancy i s t h a t t h e mounting system i n t h e 4 M x 7H is more i n t r u s i v e ( s e e Figure 12)--producing a n a c e l l e of i n c r e a s e d c r o s s - s e c t i o n a l a r e a , i s c l o s e r t o t h e p r o p e l l e r plane.

and t h e support system ..C mAcn 0.8s NACELLE ALPHA -1.

ANGLE EXPERIMENT OF OUADPAN SYMBOL ATTACK SYMBOL 1.1 1.1 1 .o 0 . 9 0.8 0.7 0.0 0 . 1 0.1 0.6 0 . 8 1.0 1.2 NOH-DWENSLONAL PROP RADIUS NON-DIMENSIONAL PROP RADIUS 0.2 0.1 0.0 > -0.1 -0.2 -0.3 0.0 o 0.. 0.6 0.8 1 . 0 1.2 "W4-DIYENSIDHAL PROP RADIUS NON-DIMENSIONAL PROP RADIUS Rake Data for Axial and Lateral Velocity Component at Mach 0 . 8 5 Figure 8 7 .

*. c MACH 0.W NACELLE ALPHA - I * ANGLE EXPERIMEHT OF QUADPAN SYMBOL ATTACK SYMBOL 0. I..

0.0 0 I 0.1 1.' 0 . 8 1.0 1.1 NOM-DIUEWIIOIIAL PROP RADIUS NOH-OIMENSIOMAL PROP RADIUS Figure 8 8 . Rake Data for Vertical Velocity Component W a t Mach 0.85 MACH = 0 . 4 BLADE ANGLE = 4 9 O

- PREDICTED, PROPVRTX

0 MEASURED, 16-FOOT

--- HAMILTON STANDARD

Figure 89. Thrust Coefficient Data at Mach 0 . 4 MACH = 0 . 4 BLADE ANGLE = 4 9 '

- PREDICTED, PROPVRTX

0 MEASURED, 16-FOOT

--- HAMILTON STANDARD

1 . 2 -

n V 0 . 8 - 0.4-

0.0 J" , I I

1.8 2 . 2 2 . 5 3 . 0 ADVANCED R A T I O , J Power Coefficient Data at Mach 0 . 4 Figure 9 0 .

MACH = 0.165, q = 40

- PREDICTED, = 3 8 '

- - - - MEASURED, f l = 38O

-e-- PREDICTED, = 4 0 ' 0.8

*......* MEASURED, f l = 4 0 '

0.6

-

V G z w

-

2 0.4 U LL W V I- v) = 0 . 2 c

-

I- 0 . 0 ADVANCE R A T I O , J Figure 9 1 . Thrust Coefficient Data at Mach 0 . 1 6 5 MACH = 0.165, q = 90

- PREDICTED, p = 3 8 O

- - - -

MEASURED, 6 = 3 8 O

- - - -

PREDICTED, p = 4 0 '

...-.-.- MEASURED, 0 = 40°

1 . 2 K 0.4 W n 0.0 ADVANCE R A T I O , J Figure 9 2 . Power Coefficient Data at Mach 0.165 When t h e t h r u s t a x i s of t h e p r o p e l l e r i s i n c l i n e d t o t h e f r e e s t r e a m , t h e p r o p e l l e r b l a d e s no l o n g e r e x p e r i e n c e uniform aerodynamic l o a d i n g but are s u b j e c t t o c y c l i c a l l y changing f o r c e s , the magnitude, of which depends on t h e b l a d e s azimuthal l o c a t i o n . The p e r i o d of t h e s e changes i s t h e same as t h a t which t h e b l a d e s t a k e t o complete one r e v o l u t i o n . Some i n s i g h t i n t o t h e o r i g i n of t h e s e l o a d s can be o b t a i n e d by examining t h e behavior of a b l a d e element d u r i n g t h e c o u r s e of r o t a t i o n . F i g u r e 93 shows s i d e and f r o n t views of a p r o p e l l e r d i s k w i t h t h e t h r u s t axis i n c l i n e d a t at t o t h e f r e e s t r e a m v e l o c i t y Vo. The f r e e s t r e a m v e l o c i t y can be resolved

i n t o component Vocos a , normal t o t h e p l a n e of r o t a t i o n , and component

V s i n a t p a r a l l e l t o t h e p l a n e of r o t a t i o n .

A b l a d e s e c t i o n o p e r a t e s w i t h an e q u i v a l e n t f r e e s t r e a m v e l o c i t y of magnitude V c o s a t and a r o t a t i o n a l component w r + VOsinar cos^, c o n s t a n t 0 t i n d i r e c t i o n but v a r y i n g i n magnitude w i t h azimuthal p o s i t i o n . Conse- q u e n t l y , t h e b l a d e can be c o n s i d e r e d t o o p e r a t e a t a v a r y i n g advance r a t i o . T h i s v a r i a t i o n i s s i n u s o i d a l and r e a c h e s a maximum a l o n g c L = 0 degrees and a minimum along fi = 180 degrees.

Consider t h e blade element shown i n F i g u r e 9 4 . The b l a d e element undergoes c y c l i c changes i n l o c a l b l a d e a n g l e of a t t a c k . T h i s v a r i a t i o n i s p l o t t e d a g a i n s t azimuthal l o c a t i o n f o r t h r e e r a d i a l s t a t i o n s of 0.35, 0.55, and 0.75 i n F i g u r e 95. Subsequently, t h i s g i v e s rise t o a c y c l i c b l a d e t h r u s t and torque. The c y c l i c t h r u s t g i v e s rise t o a s t e a d y yawing moment while t h e c y c l i c t o r q u e g i v e s rise t o a normal f o r c e . The reduc- t i o n i n forward v e l o c i t y from Vo t o V o c o s a reduces t h e advance r a t i o t ’ c o r r e s p o n d i n g l y and, as w i l l be shown l a t e r , r e s u l t s i n a J c o s a s h i f t i n t t h e performance d a t a .

F i g u r e 96 shows t h e v a r i a t i o n of s t e a d y p r o p e l l e r t h r u s t c o e f f i c i e n t w i t h t h r u s t a x i s a n g l e of a t t a c k . The o f f s e t of t h e minima of t h e s e c u r v e s i s due t o t u n n e l and model s u p p o r t upwash. As a n g l e of a t t a c k is i n c r e a s e d o r decreased from t h e s e minima ( n e t z e r o a n g l e of a t t a c k ) , t h e t h r u s t i n c r e a s e s . The s e n s i t i v i t y of t h r u s t t o a n g l e change i s a maximum a t t h e h i g h e s t advance r a t i o of 1.8 and a minimum a t an advance r a t i o of 0.9. T h i s i s d i r e c t l y r e l a t e d t o t h e e f f e c t i v e advance r a t i o J c o s a t.

F i g u r e 97 shows c u r v e s of t h e t h r u s t c o e f f i c i e n t v e r s u s advance r a t i o a t a z e r o a n g l e of a t t a c k and 16 d e g r e e s a n g l e of a t t a c k . These c u r v e s a l s o show t h e i n c r e a s e i n t h r u s t due t o t h r u s t i n c l i n a t i o n . F i g u r e 98 shows t h e two curves p l o t t e d v e r s u s J c o s a r and they almost c o l l a p s e t o a t ’ s i n g l e curve. The r e a s o n t h e y do n o t completely c o l l a p s e is t h e r e i s some r e s i d u a l upwash due t o t h e s u p p o r t system, and t h e geometric a n g l e of a t t a c k needs t o be a d j u s t e d f o r t h a t i n t e r f e r e n c e e f f e c t .

A p r o p e l l e r a t a n g l e of a t t a c k produces a normal f o r c e and an accom- panying yawing moment due t o t h e f l u c t u a t i n g p e r i o d i c l o a d s on t h e b l a d e s as e x p l a i n e d earlier. The normal f o r c e i n c r e a s e d l i n e a r l y w i t h p r o p e l l e r a n g l e of a t t a c k as shown f o r t h e 1/9-scale SR-7 a t Mach 0 . 4 o b t a i n e d i n t h e 16-Ft Transonic Aerodynamics Wind Tunnel. On i n c r e a s i n g t h e advance r a t i o , t h e normal f o r c e c o e f f i c i e n t i n c r e a s e d g i v i n g a f a m i l y of s t r a i g h t l i n e s . T h e o r e t i c a l l y , t h e y should a l l p a s s through t h e o r i g i n , because there should be no normal f o r c e a t z e r o a n g l e of a t t a c k , b u t because of t h e r e s i d u a l upwash i n t h e t u n n e l , t h e r e is some normal f o r c e a t z e r o alpha.

F i g u r e 99 shows normal f o r c e d a t a from t h e 4 M x 7M Subsonic Wind Tunnel a t Mach 0.165. Here t h e d a t a i s n o t l i n e a r b u t shows t h e v a r i a t i o n due t o changes i n upwash caused by t h e s u p p o r t system a t t h e d i f f e r e n t a n g l e s of a t t a c k .

*-

Vo - FREE STREAM VELOCITY a t - PROPELLER ANGLE OF ATTACK Figure 93. Propeller at Angle o f Attack REFERENCE VELOCITY

/ 0 I A C R A M . a = 0

t Changes in- Blade Angle of Attack Due to T h r u s t Axis Inclination Figure 94.

SR-7L 119 SCALE PROPELLER J = 1.8, p = 49", a = loo t 8.

u W n I n -I 0.

-8.

Figure 95. Cyclic Variation o f Blade Environment Due to T h r u s t Axis Inclination LOW-SPEED WIND TUNNEL D A T A MACH = 0.2 q = 60, p= 3 8 O 0 . 6 1

5 0.4

w

u

U U W P I -0.2 I I I I 1 -40 0 . 0 4.0 8 . 0 12.0 16.0 ANGLE OF ATTACK, n Figure 96. Thrust Coefficient Versus Angle o f Attack MACH = 0.165, q = 4 0 BLADE ANGLE = 38' -0- PROPELLER A T 0 ' ANGLE OF ATTACK --A-- PROPELLER A T 16' ANGLE OF A T T A C K 0.8 0 . 6 V - I-' z w

-

2 0 . 4 U U W V

5 0 . 2

LL f t - 0 . 0 ADVANCE RATIO, J T h r u s t Coefficient Versus Advance Ratio a t Two Angles Figure 97.

of Attack

*-

MACH = 0.165, q = 40 BLADE ANGLE = 3 8 O -0- PROPELLER A T Oo ANGLE OF ATTACK --A-- PROPELLER A T 1 6 O ANGLE OF ATTACK 0 . 8 1 . 2 1 . 6 2.0 ADVANCE R A T I O , JCOS a t Collapse of Thrust Coefficient Data on J Cos a t Figure 98.

MACH = 0.165 q = 4 0 , p = 40 C V 0.0 ! I I I I 1 - 4 . 0 0 . 0 4 . 0 8 . 0 12.0 16.0 ANGLE OF ATTACK, a t Figure 9 9 . Normal Force Coefficient Data Versus Angle of Attack 8.0 CONCLUDING REMARKS 1. Performance d a t a from t h e s e wind t u n n e l tests were v a l i d a t e d i n two ways. The d a t a f o r t h e G I 1 c o n f i g u r a t i o n were v a l i d a t e d by comparison w i t h published d a t a f o r t h e G I 1 a i r c r a f t . Good agreement was found.

Data f o r both t h e G I 1 and t h e PTA model c o n f i g u r a t i o n were compared w i t h a n a l y t i c a l p r e d i c t i o n s made w i t h t h e Lockheed codes QUADPAN and PROPVRTX. Again, t h e agreement was good.

2. A t moderate a n g l e s of a t t a c k and subsonic speeds, t h e impact of t h e PTA m o d i f i c a t i o n s on a i r c r a f t performance were t o i n c r e a s e l i f t s l i g h t l y and i n c r e a s e d r a g by about 15 percent. The d r a g increment i n c r e a s e d r a p i d l y above Mach 0 . 6 , however, and a t t h e d e s i g n c r u i s e Mach number of 0.8, t h e d r a g increment w a s about 30 p e r c e n t of t h e b a s i c G I 1 drag. Maximum l i f t c o e f f i c i e n t s were reduced by t h e PTA mod i f i c a t ions.

3. The PTA m o d i f i c a t i o n s reduced l o n g i t u d i n a l s t a b i l i t y s l i g h t l y but had n e g l i g i b l e e f f e c t s on l a t e r a l and d i r e c t i o n a l s t a b i l i t y .

4 . R o l l c o n t r o l power was s i g n i f i c a n t l y reduced from t h e G I 1 values-- p r i m a r i l y because some of t h e G I 1 s p o i l e r p a n e l s were d e a c t i v a t e d f o r t h e PTA i n s t a l l a t i o n . Rudder and e l e v a t o r e f f e c t i v e n e s s , on t h e o t h e r hand, were only s l i g h t l y a f f e c t e d .

5. Test r e s u l t s showed t h a t a l e a d i n g edge e x t e n s i o n (LEX) designed f o r t h e wing on t h e inboard s i d e of t h e n a c e l l e would l i k e l y reduce a i r - c r a f t d r a g i n t h e high-speed d r a g rise range i f it were needed.

6. Three-component v e l o c i t y surveys i n t h e high-speed flow f i e l d j u s t behind t h e prop plane showed reasonable c o r r e l a t i o n w i t h t h e o r y i n some cases, and n o t s o good c o r r e l a t i o n i n o t h e r s . C o r r e l a t i o n was b e s t f o r t h e t r a n s v e r s e components where measurement accuracy was b e t t e r . A major o b j e c t i v e of t h e s e tests w a s t o determine i f t h e r e was a s i g n i f i c a n t d e t e r i o r a t i o n of t h e c o r r e l a t i o n a t t h e t r a n s o n i c speeds. The d a t a i n d i c a t e t h a t t h i s w a s n o t t h e case.

7. A concern a t t h e beginning of t h i s program had t o do w i t h t h e d e s i g n of small-scale propfan r o t o r s t o p r o p e r l y s i m u l a t e t h e s i g n i f i c a n t c h a r a c t e r i s t i c s of t h e f u l l - s c a l e r o t o r s . I s o l a t e d propfan r o t o r tests were performed t o provide d a t a so t h a t scale c o r r e c t i o n s could be applied. R e s u l t s showed a good c o r r e l a t i o n between p r e d i c t e d and measured rotor performance.

OT

, APPENDIX A DESIGN OF THE LEX (LEADING EDGE EXTENSION) Transonic d r a g of t h e PTA c o n f i g u r a t i o n w a s a major concern i n t h e d e s i g n program. It was recognized t h a t t h e combination o f : (a) wing PTA n a c e l l e , ( b ) local-wing-angle-of-attack unsweeping by a d d i t i o n of t h e i n c r e a s e s due t o t h e s w i r l of t h e s l i p s t r e a m , and (c) flow c h a n n e l i n g by t h e f u s e l a g e / w i n g / n a c e l l e c o n f i g u r a t i o n h e l d t h e p o t e n t i a l f o r s i g n i f i c a n t u n p r e d i c t a b l e d r a g rise.

I n r e c o g n i t i o n of t h e s e concerns, s e v e r a l d r a g r e d u c t i o n d e v i c e s were explored a n a l y t i c a l l y , and one w a s s e l e c t e d t o be t e s t e d i n t h e t u n n e l program. The l a t t e r was a l e a d i n g edge e x t e n s i o n (LEX) on t h e inboard s i d e of t h e n a c e l l e . Its purpose w a s t o recamber t h e wing i n that r e g i o n most l i k e l y t o be n e g a t i v e l y a f f e c t e d by s l i p s t r e a m s w i r l .

DESIGN OBJECTIVES The LEX d e s i g n employed a d e s i r e d o r t a r g e t p r e s s u r e d i s t r i b u t i o n a s t h e "optimization" c r i t e r i o n because s u r f a c e p r e s s u r e s may be used t o q u a l i t a t i v e l y e v a l u a t e t h e d r a g c h a r a c t e r i s t i c s of a c o n f i g u r a t i o n . I n t h e t r a n s o n i c regime, t h e y can r e v e a l t h e presence of shock waves and t h u s t h e wave d r a g a s s o c i a t e d w i t h t h e c o n f i g u r a t i o n . Design e x p e r i e n c e allows one t o d e r i v e a s u i t a b l e t a r g e t p r e s s u r e d i s t r i b u t i o n f o r a g i v e n s u r f a c e .

I n t h e t r a n s o n i c c a s e , t h e target wing s e c t i o n a l p r e s s u r e on t h e upper s u r f a c e should have no l e a d i n g edge peaks which l e a d t o premature shock wave compressions, b u t r a t h e r a p r e s s u r e c o e f f i c i e n t d i s t r i b u t i o n t h a t is well-rounded, i n d i c a t i n g a c o n t r o l l e d flow a c c e l e r a t i o n t o s u p e r s o n i c v e l o c i t i e s , followed by a r e t u r n t o s u b s o n i c f l o w through a n e a r s h o c k l e s s recompression. The a d v e r s e p r e s s u r e g r a d i e n t approaching t h e t r a i l i n g edge should be g e n t l e enough s o t h a t no r e a r s e p a r a t i o n occurs.

C o l l e c t i v e l y , t h e s e c t i o n p r e s s u r e d i s t r i b u t i o n s can be a s s e s s e d by examining t h e wing i s o b a r p a t t e r n . Any l o s s of i s o b a r sweep due t o l o c a l i z e d h i g h s u c t i o n peaks and s t e e p a d v e r s e p r e s s u r e g r a d i e n t s should be avoided. S i n c e t h e o r i g i n a l G I 1 wing embodied t h e aforementioned d e s i g n c r i t e r i o n , t h e t a r g e t adopted i n t h e LEX d e s i g n was t o r e s t o r e t h e PTA wing flow t o p a t t e r n s similar t o t h a t of t h e G I 1 wing.

DESIGN PROCEDURES The planform e x t e n t of t h e m o d i f i c a t i o n w a s e s t a b l i s h e d as shown i n F i g u r e A - 1 , and t h r e e wing c o n t r o l s e c t i o n s were chosen. A t t h e s e t h r e e s e c t i o n s t h e wing chord was extended forward and reshaped t o b e t t e r accommodate t h e upwash i n t h e j u n c t u r e area from t h e s l i p s t r e a m s w i r l .

The e x t e n s i o n was blended w i t h t h e n a c e l l e , care being t a k e n t o keep t h e f r o n t a l area a minimum.

The b a s i c d e s i g n procedure i s shown s c h e m a t i c a l l y i n F i g u r e A-2.

F i r s t , SUNTAN, a two-dimensional t r a n s o n i c code, was employed t o produce wing s e c t i o n s approaching t h e d e s i r e d c o n t o u r s a t t h e c o n t r o l s e c t i o n s .

The upper s u r f a c e i s more c r i t i c a l t h a n the lower and w a s blended i n t o t h e o r i g i n a l wing s o as n o t t o cause premature recompression o r "kinks" i n t h e p r e s s u r e d i s t r i b u t i o n . On t h e lower s u r f a c e , t h e j u n c t i o n area w i l l cause a sudden p r e s s u r e drop, b u t s i n c e t h e lower s u r f a c e is less c r i t i c a l t h i s should not s i g n i f i c a n t l y a f f e c t t h e performance of t h e LEX. A s e c t i o n r e d e s i g n i s shown i n F i g u r e A-3. The three-dimensional LEX r e g i o n was t h e n b u i l t by a p p r o p r i a t e l o f t i n g between t h e s e c o n t r o l s e c t i o n s u s i n g CATZA.

The design procedure is used a second t i m e u s i n g a three-dimensional t r a n s o n i c code--a modified v e r s i o n of t h e Jameson f u l l - p o t e n t i a l flow code, FL022NM, which i n c l u d e s n a c e l l e and s l i p s t r e a m e f f e c t s . Once a s a t i s f a c t o r y wing s u r f a c e was o b t a i n e d , t h e QUADPAN program was used t o contour t h e LEX/nacelle j u n c t i o n region. A s might be expected, some of t h e s t e p s i n t h i s p r o c e s s were repeated t o r e s o l v e any undesired behavior.

::1

Figure A-1. Planform View of Leading Edge Extension (LEX) PROPVRTX PROPELLER COMPARE OUTPUT TO DESIRED PRESSURE ACCEPTABLE REFINE DESIGN USING NEXT CODE Figure A-2. Aerodynamic Design Procedure for LEX

LEX GEOMETRY - SECTION A T = 0 . 3 0 8

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I I Figure A-3. Airfoil Section Modification for LEX

APPENDIX B CALIBRATION AND D A T A REDUCTION FOR 5-HOLE PROBES The 5-hole probes were c a l i b r a t e d i n t h e Lockheed-Georgia Company C F W T Transonic Wind Tunnel a t Mach numbers t o 0 . 9 5 . The c a l i b r a t i o n c o n s i s t e d of a procedure of p l a c i n g e a c h r a k e , i n t u r n , i n known flow c o n d i t i o n s and measuring t h e p r e s s u r e a t each o r i f i c e on t h e rake. The flow c o n d i t i o n s were v a r i e d by changing t h e roll and p i t c h a n g l e s e t t i n g s of t h e r a k e and v a r y i n g t h e test s e c t i o n Mach number. The v a r i a t i o n s i n flow c o n d i t i o n s are shown below: P i t c h Angle ( T h e t a ) R o l l Angle ( P h i ) Mach Number 0 t o 22 Degrees i n 0 t o 360 Degrees i n 0.3, 0.6, and 0.9 2-Degree Increments 15-Degree Increments The a c q u i s i t i o n of t h e s e d a t a r e q u i r e d 900 runs w i t h e a c h run con- s i s t i n g of a s i n g l e flow c o n d i t i o n . A d d i t i o n a l l y , d a t a were acquired f o r Mach numbers from 0.2 t o 0.90 i n M = 0.1 increments and, i n t h e range from 0.90 t o 0.95 i n M = 0.05 increments f o r p i t c h a n g l e s of 0, 10, and 20 d e g r e e s and r o l l a n g l e s of 0 and 90 degrees.

The f i r s t r a k e was a p r o t o t y p e and was not f u l l y c a l i b r a t e d . It was t e s t e d f o r development of t h e c a l i b r a t i o n procedure and checkout of t h e rake design. Rakes 2 and 3 were f u l l y c a l i b r a t e d i n test CFWT 101 and t e s t CFWT 104, r e s p e c t i v e l y .

The p r e s s u r e d a t a measured a t e a c h o r i f i c e on a probe was used t o c a l c u l a t e f u n c t i o n s f o r probe-sensed flow p i t c h a n g l e ( a l p h a ) , yaw a n g l e ( p s i ) , t o t a l pressure ( H ) , and dynamic p r e s s u r e i n t h e f o l l o w i n g manner:

where C p is t h e minimum f ( q ) = cp5 - c p m

m of c p 1’ C p 2 7 C p y and c p 4 The d a t a from t h e cone s t a t i c were not used i n g e n e r a t i n g t h e cali- b r a t i o n due t o d i f f i c u l t i e s i n o b t a i n i n g c o n s i s t e n t d a t a from them.

The c a l i b r a t i o n procedure f i r s t produced a set of f(a1pha) and f ( p s i ) f u n c t i o n s f o r c o n s t a n t probe p i t c h and yaw angles. These f u n c t i o n s , through i n t e r p o l a t i o n , were used t o determine maps of a l p h a v e r s u s p s i f o r c o n s t a n t f(a1pha) and f ( p s i ) . T h i s map provided a d i r e c t look-up t a b l e f o r a l p h a and p s i once t h e a n g l e f u n c t i o n s were c a l c u l a t e d .

The f ( H ) and f ( Q ) p r e s s u r e f u n c t i o n s were a l s o i n t e r p o l a t e d and s t o r e d i n look-up t a b l e s a g a i n s t c o n s t a n t v a l u e s of f(a1pha) and f ( p s i ) .

T h i s u n i q u e l y i d e n t i f i e d t h e p r e s s u r e f u n c t i o n s over t h e c a l i b r a t e d a n g l e range .

I n t h e c u r r e n t computer s o f t w a r e implementation of t h e probe c a l i b r a - t i o n procedure, t h e range of f ( a 1 p h a ) and f ( p s i ) w a s set t o c o n s i s t e n t v a l u e s f o r a l l c a l i b r a t e d probes. T h i s r e s u l t e d i n t h e c a l i b r a t e d range of a l p h a and p s i f o r e a c h probe t o be s l i g h t l y d i f f e r e n t from t h e o t h e r probes. The v a r i a t i o n from probe t o probe was caused by small d i f f e r e n c e s i n t h e probe t i p s and p o s s i b l y c u r v a t u r e of the probe s h a f t s . The range of f(a1pha) and f ( p s i ) w a s determined by t h e performance of t h e worst probe of t h e group of probes t h a t a c a l i b r a t i o n was t o be prepared f o r .

T h i s l i m i t a t i o n could be overcome by s t o r i n g more d e t a i l e d c a l i b r a t i o n i n f o r m a t i o n about each probe.

s t e p s taken i n u s i n g t h e c a l i b r a t i o n s t o determine t h e unknown The l o c a l f l o w c o n d i t i o n s from t h e known f i v e probe p r e s s u r e s d u r i n g a flow t e s t were as f o l l o w s : survey 1. C a l c u l a t e f ( a 1 p h a ) and f(psi) as shown above.

2. Perform double i n t e r p o l a t i o n u s i n g a l p h a v a l u e s s t o r e d v e r s u s f(a1pha) and f ( p s i ) t o g e t flow p i t c h angle.

3. Perform double i n t e r p o l a t i o n u s i n g p s i v a l u e s s t o r e d v e r s u s f ( a 1 p h a ) and f ( p s i ) t o get flow yaw angle.

4. Perform double i n t e r p o l a t i o n u s i n g f(H) v a l u e s s t o r e d v e r s u s f(a1pha) and f ( p s i ) t o g e t f(H) value.

5. Perform double i n t e r p o l a t i o n u s i n g f ( q ) v a l u e s s t o r e d v e r s u s f(a1pha) and f ( p s i ) t o g e t f(q) value.

6. C a l c u l a t e l o c a l dynamic pressure (4) using:

q = f ( q ) * qo

7. C a l c u l a t e l o c a l t o t a l p r e s s u r e (H) u s i n g :

H = f ( H ) * qo + po

8. C a l c u l a t e l o c a l v e l o c i t y v e c t o r components using:

U = ( f ( q ) / (1.0 + tan(alpha)**2 + t a n (psi)**2) )**0.5

V = U * t a n ( p s i )

W = U * tan(a1pha)

A l l d a t a are i n - t h e probe axis system which can t h e n be t r a n s - formed i n t o any d e s i r e d axis system.

A f t e r t h e above c a l c u l a t i o n s were performed, t h e b a s i c flow param- eters a t each survey p o i n t were known.

APPENDIX C ANALYTICAL PREDICTIONS Developments i n computational aerodynamics i n t h e p a s t 15 y e a r s have p e r m i t t e d r a p i d , e f f i c i e n t , and r e l a t i v e l y a c c u r a t e numerical s o l u t i o n s of complex subsonic and t r a n s o n i c flows, The complex aerodynamic flow- f i e l d s t h a t p r e s e n t computational methods are c a p a b l e of a n a l y z i n g are e x e m p l i f i e d by numerical s i m u l a t i o n s f o r wing-mounted t r a c t o r turboprop The i n s t a l l a t i o n s w i t h i n t e g r a t e d n a c e l l e s and s w i r l i n g s l i p s t r e a m s .

following paragraphs d e s c r i b e t h e computational methods which were used t o assess t h e PTA c o n f i g u r a t i o n , i d e n t i f y p o t e n t i a l problem areas, assess p o s s i b l e d e s i g n m o d i f i c a t i o n remedies and d e s i g n t h e LEX region.

PROPELLER/PANEL METHOD THEORY Two primary computer programs are u t i l i z e d . The f i r s t i s a r e l i a b l e p r o p e l l e r performance c o d e , (PROPVRTX), which a l l o w s major c h a r a c t e r i s t i c s of t h e s l i p s t r e a m t o be determined. The code p r e d i c t s t h e s l i p s t r e a m p r o p e r t i e s produced by a p r o p e l l e r of g i v e n geometry and advance r a t i o .

The u n d i s t u r b e d s l i p s t r e a m i s d e f i n e d by i t s boundary shape and t h e r a d i a l d i s t r i b u t i o n of t h e l o c a l v o r t i c i t y , induced v e l o c i t y , and s w i r l angle.

are i n v e r y good agreement w i t h test d a t a . The The r e s u l t s from t h i s code second primary code i s QUADPAN, Lockheed's advanced low-order, t h r e e - dimensional panel method. QUADPAN has been widely used i n t h e a i r p l a n e d e s i g n p r o c e s s f o r t h e a n a l y s i s of g e o m e t r i c a l l y complex c o n f i g u r a t i o n s and t h e r e s u l t s have been h i g h l y r e l i a b l e . For t h e PTA c o n f i g u r a t i o n , QUADPAN is uniquely q u a l i f i e d t o provide a good geometric r e s o l u t i o n i n t h e i n t r i c a t e n a c e l l e / i n l e t region. A d d i t i o n a l l y , t h e method i s r e l a t i v e l y i n s e n s i t i v e t o t h e s i z e and shape of t h e panel elements because it employs an i n t e r n a l p o t e n t i a l ( D i r i c h l e t ) boundary c o n d i t i o n . The s o l u t i o n o u t p u t c o n s i s t s of l o c a l p r e s s u r e c o e f f i c i e n t s f o r a l l s u r f a c e p a n e l s , flow c o n d i t i o n s a t s p e c i f i e d p o i n t s i n t h e e x t e r n a l flow f i e l d , and f o r c e s and moments f o r t h e t o t a l c o n f i g u r a t i o n as w e l l as f o r i t s component p a r t s . The two codes are i n t e r f a c e d w i t h t h e e f f e c t of t h e slipstream being simulated through a r e s t a t e m e n t of t h e s u r f a c e boundary c o n d i t i o n s t o i n c l u d e v e l o c i t y p e r t u r b a t i o n s due t o t h e p r o p e l l e r . The p r e s s u r e on t h e s u r f a c e s washed by t h e s l i p s t r e a m are a p p r o p r i a t e l y c o r r e c t e d f o r t h e p r e s s u r e i n c r e a s e a c r o s s t h e p r o p e l l e r d i s k . T h i s approach i s similar t o one p r e v i o u s l y employed on a wing a n a l y s i s program w i t h s a t i s f a c t o r y r e s u l t s .

PROPELLER EFFECTS SUBROUTINE The e f f e c t of t h e p r o p e l l e r on t h e n a c e l l e i s c a l c u l a t e d by means of a r i g i d s l i p s t r e a m model which provides t h e v a r i a t i o n of a x i a l v e l o c i t y a s e m i - i n f i n i t e h e l i c a l v o r t e x system. The s l i p s t r e a m c o n t r a c t i o n i s w i t h found from t h e a p p l i c a t i o n of t h e c o n t i n u i t y p r i n c i p l e t o t h e s l i p s t r e a m .

A s t h e r e i s a p r e s s u r e rise through t h e p r o p e l l e r d i s k due t o t h e addi- t i o n a l energy i n p u t from t h e p r o p e l l e r , i t i s n e c e s s a r y t o i n c l u d e t h i s e x p l i c i t l y i n t h e c a l c u l a t i o n . The p r e s s u r e 7 ;e can be expressed i n terms of t h e v e l o c i t y of t h e f u l l y c o n t r a c t e d s l i p s t r e a m . The proper l e v e l of t h r u s t is s i m u l a t e d by s c a l i n g t h e a x i a l and t a n g e n t i a l v e l o c i t y d i s t r i b u t i o n s t o correspond t o a d e s i r e d power c o e f f i c i e n t Cp, and advance r a t i o J. The s l i p s t r e a m d a t a were d e r i v e d from experimental rake d a t a f o r a g i v e n ?ewer c o e f f i c i e n t and advance r a t i o and i n t e g r a t e d t o g i v e an average s w i r l a n g l e and average a x i a l v e l o c i t y increment. The r a d i a l d i s t r i b u t i o n i s k e p t t h e same f o r a l l advance r a t i o s and t h e magni- t u d e is c o r r e c t e d by s c a l i n g t h e i n t e g r a t e d v a l u e s t o g i v e t h e d e s i r e d t h r u s t l e v e l .

COMPUTATIONAL MODELING The b a s i c s t e p s i n computational aerodynamic d e s i g n are: ( I ) t h e s u r f a c e geometric d e f i n i t i o n , ( 2 ) aerodynamic a n a l y s i s , and ( 3 ) perform- ance e v a l u a t i o n . The geometry must be i n a form t h a t allows r a p i d access, and i t must d e f i n e t h e c o n f i g u r a t i o n t o t h e a p p r o p r i a t e d e g r e e of accuracy and d e t a i l .

The e n g i n e e r i n g a n a l y s i s is c a r r i e d o u t by e i t h e r tests o r by a computation of a mathematical simulation. C e r t a i n requirements f o r t h e computational codes must be m e t t o e n s u r e t h e i r e f f e c t i v e n e s s i n t h e d e s i g n procedure. The numerical model must a d e q u a t e l y s i m u l a t e t h e p e r t i n e n t physics of t h e flow, such a s t h e i n f l u e n c e of t h e p r o p e l l e r s l i p s t r e a m and of t h e presence of shock waves i n t h e t r a n s o n i c regime.

The computer codes should be capable of handling complex geometric shapes such as t h e i n t r i c a t e n a c e l l e - i n l e t r e g i o n t y p i c a l of a t r a c t o r turboprop c o n f i g u r a t i o n . The codes a r e a l s o r e q u i r e d t o have a r a p i d turn-around t i m e , and they must n o t depend on e x t e n s i v e manipulation of t h e i n p u t d a t a t o a c h i e v e a s u c c e s s f u l run. The i n p u t and o u t p u t d a t a management should p r e f e r a b l y be automated. Performance e v a l u a t i o n is c a r r i e d o u t u s i n g t h e a c q u i r e d d a t a t o d e t e r m i n e t h e e x t e n t t o which t h e d e s i g n o b j e c t i v e s have been m e t and t o d e c i d e what c o n f i g u r a t i o n a l changes are needed. To c a r r y o u t t h e e v a l u a t i o n r a p i d l y , automated p l o t t i n g r o u t i n e s of t h e important parameters must be a v a i l a b l e and where t h e c a p a b i l i t y e x i s t s p l o t t i n g c o n t o u r s and s u r f a c e c o l o r shading i s a b i g h e l p i n examining t h e enormous amount of i n f o r m a t r o n t h a t can be e x t r a c t e d from t h e s e computational codes.

S u r f a c e Geometric D e f i n i t i o n The C A T I A computer system, which was developed by D a s s a u l t of France, w a s used t o manipulate and s t o r e t h e c o n f i g u r a t i o n geometry. C A T I A , is an i n t e r a c t i v e , three-dimensional geometry system f o r computer-aided d e s i g n and manufacturing. The system a l l o w s t h e r a p i d g e n e r a t i o n of e i t h e r t h e complete c o n f i g u r a t i o n o r any p a r t of t h e c o n f i g u r a t i o n , and allows t h e p a r t t o be viewed as a three-dimensional drawing from any view p o i n t . A two-dimensional s e c t i o n a t any p l a n e can a l s o be defined. I n a d d i t i o n , t h e geometric d e f i n i t i o n can be o b t a i n e d i n any d e s i r e d c o o r d i n a t e system.

These f e a t u r e s of t h e CATIA system p e r m i t r a p i d and s i m p l i f i e d s u c c e s s i v e c o n f i g u r a t i o n changes.

Engineering Analysis The QUADPAN program meets t h e code requirements s t i p u l a t e d above with t h e e x c e p t i o n of c a l c u l a t i n g t r a n s o n i c flows w i t h shock waves. QUADPAN allows adequate r e p r e s e n t a t i o n of t h e complex i n l e t geometry by subdivid- i n g t h e s u r f a c e i n t o a l a r g e number of small panels. Unfortunately, run t i m e and run c o s t i n c r e a s e almost w i t h t h e t h i r d power of t h e number of p a n e l s , so compromises must be made. A f u l l model w i t h 5,000 p a n e l s t y p i c a l l y took 20 t o 30 minutes of CPU t i m e and 500 megabytes of high-speed d i s k s t o r a g e on t h e Lockheed CRAY W / 2 4 d u a l processor supercomputer and is c u r r e n t l y seen as t h e l i m i t of modeling d e n s i t y .

I n o r d e r t o properly s i m u l a t e t h e physics of t h e flow, t h r e e flow s i t u a t i o n s r e q u i r e s p e c i a l a t t e n t i o n . These are flow from t h e wing trail- i n g edge, t h e j e t from t h e engine e x h a u s t , and t h e i n t e r n a l flow through t h e engine.

A i r c r a f t Wake Treatment I n t h e r e a l world, a f l u i d l e a v e s from t h e s h a r p t r a i l i n g edge of a wing as a r e s u l t of v i s c o u s e f f e c t s , c r e a t i n g t h e so-called Kutta con- d i t i o n , where t h e outflow i s smooth and n e a r l y tangent t o t h e t r a i l i n g edge. Since a panel program does n o t model v i s c o s i t y , it i s necessary t o i n t r o d u c e a t h i n v o r t e x s h e e t a t t h e s e p a r a t i o n l i n e along t h e t r a i l i n g edge. This s h e e t of v o r t i c i t y or wake l o c a t e s t h e l e a d i n g edge stagna- t i o n p o i n t , and so f i x e s t h e wing's c i r c u l a t i o n , or l i f t . C l e a r l y , t h e proper treatment of t h e wake is of importance i n a method t h a t assumes t h e flow is otherwise i r r o t a t i o n a l and i n v i s c i d .

The s i t u a t i o n f o r t h e PTA c o n f i g u r a t i o n i s complicated by t h e l a r g e over-the-wing n a c e l l e , s i n c e , i n r e a l i t y , a Kutta c o n d i t i o n i s a l s o s a t i s f i e d along t h e s h a r p edges of t h e n a c e l l e ' s a f t end, and around t h e circumference of t h e t u r b o s h a f t e n g i n e ' s exhaust nozzle. To model t h i s geometry, i t w a s decided t o approximate t h e a f t end of t h e n a c e l l e a s a cone t a p e r i n g down t o a v e r y long c y l i n d e r r e p r e s e n t i n g t h e engine exhaust plume. Although both t h e exhaust cone and exhaust plume a r e r e p r e s e n t e d i n t h e model as s o l i d s u r f a c e s , t h e aerodynamic f o r c e s on them a r e not d i r e c t l y included a s f o r c e s on t h e a i r c r a f t . The wing wake is then a t t a c h e d t o t h e taper of t h e exhaust cone i n a fashion s i m i l a r t o t h e treatment of t h e wing/fuselage j u n c t u r e , s e a l i n g u l t i m a t e l y along t h e s i d e s of t h e ( s o l i d ) plume.

Since a t h e o r e t i c a l l y i l l - p o s e d problem can result from over- s p e c i f y i n g s u r f a c e boundary c o n d i t i o n s , t h e exhaust plume t e r m i n a t e s i n a d i s k where only t h e p o t e n t i a l boundary c o n d i t i o n i s applied. The doublet s t r e n g t h i s t h u s set, and t h e required s o u r c e s t r e n g t h f o r t h i s c o n d i t i o n may be c a l c u l a t e d .

I n t e r n a l Flow Simulation . It i s most important t o be a b l e t o s i m u l a t e flow i n t o t h e i n l e t , and i t should correspond t o a p r e s c r i b e d mass flow r a t i o . Panel programs provide a n a t u r a l means of modeling i n l e t flow i n g e s t i o n by a s i m p l e r e s t a t e m e n t of t h e panel boundary condition. The QUADPAN code has t h e o p t i o n t o use permeable p a n e l s on which t h e normal v e l o c i t y is s p e c i f i e d by t h e u s e r t o correspond t o t h e r e q u i r e d mass flow a t t h e e n g i n e cam- p r e s s o r face. For incompressible flow, t h e v e l o c i t y induced by t h e s i n g u l a r i t i e s a t t h e c o n t r o l p o i n t of t h e permeable p a n e l w i l l e x a c t l y e q u a l t h e s p e c i f i e d v e l o c i t y . This is n o t t r u e f o r compressible flow, as t h e c o m p r e s s i b i l i t y c o r r e c t i o n i n QUADPAN assumes t h a t t h e p a n e l s are n e a r l y p a r a l l e l t o t h e f r e e s t r e a m , and t h e l i n e a r i z a t i o n of the Prandtl- G l a u e r t e q u a t i o n s upon which t h e panel method i s founded is a c c u r a t e o n l y t o f i r s t order. T h e r e f o r e , a p a n e l t h a t s t a n d s normal t o f r e e s t r e a m flow, as i n t h e case of an i n l e t f a c e o r s t a g n a t i o n p o i n t , v i o l a t e s t h e s e assumptions and w i l l have c o n s i d e r a b l e leakage of f l u i d through its s u r f a c e . Consequently, a s p e c i f i e d v e l o c i t y corresponding t o a h i g h e r mass f l o w than t h a t d e s i r e d is i n p u t f o r compressible cases. The primary powerplant t u r b o j e t e n g i n e n a c e l l e s are much easier t o s i m u l a t e and are t r e a t e d as simple flow-through d u c t s . Power e f f e c t s can be s i m u l a t e d by p l a c i n g a d i s k w i t h a p r e s c r i b e d normal v e l o c i t y a t t h e e n g i n e compressor f a c e , and t r e a t i n g t h e e x h a u s t i n t h e same way a s t h e a f t end of t h e t u r b o s h a f t engine e x h a u s t plume, w i t h a c y l i n d r i c a l wake t o e n f o r c e t h e Kutta condition.

Performance E v a l u a t i o n The c o n f i g u r a t i o n performance i s e v a l u a t e d by f i r s t d e t e r m i n i n g t h e e x t e n t t o which t h e d e s i g n o b j e c t i v e s have been m e t , and then i d e n t i f y i n g regions where aerodynamic behavior is unacceptable o r improvements are d e s i r a b l e . To make t h e e v a l u a t i o n r a p i d l y , automated p l o t t i n g f o r t h e p e r t i n e n t parameters is a v a i l a b l e . The mathematical s i m u l a t i o n r e s u l t s can be d i s p l a y e d i n f o u r d i f f e r e n t ways: ( 1 ) p r e s s u r e d i s t r i b u t i o n p l o t s , ( 2 ) v e l o c i t y v e c t o r p l o t s , ( 3 ) f o r c e and moment p l o t s , and ( 4 ) i s o - parameter p l o t s . P r e s s u r e p l o t s h e l p l o c a t e adverse p r e s s u r e g r a d i e n t r e g i o n s , as i s t h e case due t o t h e upwash from t h e propfan s l i p s t r e a m .

Velocity-vector p l o t s a l l o w f l o w f i e l d d a t a a t t h e p r o p e l l e r p l a n e t o b e examined. Force and moment p l o t s are e s s e n t i a l t o determine t h e l o a d s on v a r i o u s components of t h e c o n f i g u r a t i o n . Iso-parameter p l o t s h e l p l o c a t e r e g i o n s of unusual a i r f l o w a c t i v i t y i n areas such as t h e n a c e l l e / w i n g j u n c t u r e , and a l l o w s an assessment of t h e blending t o be made. S u r f a c e c o l o r shadings t h a t r e p r e s e n t l o c a l p r e s s u r e c o e f f i c i e n t s a r e a b i g h e l p i n examining t h e enormous amount of i n f o r m a t i o n t h a t can be e x t r a c t e d from t h e s e compu t a t iona 1 code s.

APPENDIX I ) .

DRAG D A T A ANALYSIS TURBOPROP DRAG BOOKKEEPING The p r e d i c t i o n of f u l l scale turboprop a i r c r a f t performance from wind t u n n e l d a t a r e q u i r e s a p p r o p r i a t e tests t o i d e n t i f y t h e e f f e c t of i n s t a l l a - t i o n on p r o p e l l e r performance and l i k e w i s e t h e e f f e c t of t h e s l i p s t r e a m on a i r f r a m e performance. I s o l a t e d p r o p e l l e r performance i s a p r e r e q u i s i t e t o t h e p r o p e r i d e n t i f i c a t i o n of t h e v a r i o u s t h r u s t and d r a g components such t h a t manufacturers a c t i n g w i t h i s o l a t e d e n g i n e , p r o p e l l e r , and a i r c r a f t d a t a can p r e d i c t t h e whole a i r c r a f t performance. I n t h e PTA Program, i s o l a t e d p r o p u l s i o n c h a r a c t e r i s t i c s were t o be o b t a i n e d by c a l i b r a t i o n of t h e wind t u n n e l p r o p u l s i o n hardware, i.e., t h e p r o p e l l e r and exhaust n o z z l e p r i o r t o complete model tests.

T h r u s t g e n e r a t e d by t h e i n s t a l l e d p r o p e l l e r i s t h e n compared t o t h a t of t h e i s o l a t e d p r o p e l l e r a t t h e same f r e e s t r e a m Mach number, advance r a t i o , and blade p i t c h s e t t i n g . P r o p e l l e r performance is extremely s e n s i - t i v e t o b l a d e p i t c h s e t t i n g , and because it i s s o d i f f i c u l t t o a c c u r a t e l y r e p e a t p i t c h s e t t i n g for t h e i s o l a t e d p r o p e l l e r and t h e i n s t a l l e d config- u r a t i o n , t h r e e propfan r o t o r sets were employed w i t h d i f f e r e n t blade-pitch a n g l e s f i x e d and u n d i s t u r b e d throughout t h e tests.

The d r a g p o l a r s f o r each s p e c i f i c c o n f i g u r a t i o n i n c l u d e i n t e r f e r e n c e e f f e c t s due t o t h e i n s t a l l a t i o n of t h e p r o p u l s i o n u n i t . The e f f e c t i v e d r a g c o e f f i c i e n t i s d e f i n e d as: c = c - c - c D~~~ D~~~ tISO t J E T where t h e i s o l a t e d p r o p e l l e r t h r u s t , C t , i s found from t h e i s o l a t e d p r o p e l l e r performance t e s t a t t h e same f r e e s t r e a m Mach number, advance i s t h e t h r u s t due t o t h e exhaust r a t i o , and b l a d e p i t c h s e t t i n g .

ct and i s o b t a i n e d from t h e exhaust nozzle c a l i b r a t i o n .

P r o p e l l e r i n t e r f e r e n c e d r a g i s o b t a i n e d f o r t h e g i v e n c o n f i g u r a t i o n by s u b t r a c t i n g t h e p r o p e l l e r - o f f d r a g from t h e e f f e c t i v e drag: PRESENT W I N D TUNNEL DRAG DATA ANALYSIS During t h e PTA wind-tunnel t e s t i n g , dynamic v i b r a t i o n s w i t h t h e i s o l a t e d p r o p e l l e r l n a c e l l e c o n f i g u r a t i o n precluded t h e c a l i b r a t i o n of t h e 119-scale p r o p e l l e r a t Mach numbers above 0.4; t h u s , t h e e f f e c t i v e d r a g c o e f f i c i e n t cannot be obtained. The wind t u n n e l d a t a i s reduced u s i n g t h e hub-balance t h r u s t measurement t o o b t a i n t h e model p r o p e l l e r - o n drag.

T h i s i s d e f i n e d as follows:

C = c - c

- ct

JET DPo D~~~ t~~ The propeller-on d r a g is t h e d r a g of t h e c o n f i g u r a t i o n with t h e pro- p e l l e r t h r u s t ( C t H B > , as measured by t h e hub-balance, and t h e j e t t h r u s t CDPO has a l l t h e e f f e c t s of t h e s l i p s t r e a m on t h e wing (CtJET) removed.

which i n c l u d e s a t h r u s t term r e s u l t i n g from the d e n o t a t i o n of the s l i p - stream and scrubbing d r a g due t o t h e a d d i t i o n a l dynamic p r e s s u r e w i t h i n t h e slipstream. A t h i g h Mach numbers, the s l i p s t r e a m alters t h e shock wave p a t t e r n on t h e wing and t h e r e s u l t a n t increment i n f o r c e is dependent on t h e mutual t r a n s o n i c i n t e r a c t i o n s which may o r may n o t be f a v o r a b l e .

I n t h e r e d u c t i o n of t h e t h r u s t / d r a g d a t a f o r t h e PTA 1/9-scale model, a number of d a t a s o u r c e s were u t i l i z e d . These included t h e c o r r e - l a t i o n of t h e b a l a n c e t e m p e r a t u r e c o r r e c t i o n , i n t e r n a l d r a g s f o r t h e Spey and powered n a c e l l e s , propfan d r i v e motor horsepower c o r r e l a t i o n , unpowered model d r a g p o l a r , and f i n a l l y , f o r c e components f o r t h e f u l l - u p PTA a i r p l a n e model running under power. A d d i t i o n a l components, such as tares and base d r a g components, were i n c l u d e d where a p p r o p r i a t e .

INTERNAL DRAG, SPEY NACELLES The Spey n a c e l l e s f o r t h e PTA model were instrumented more than i s normally done f o r flow-through n a c e l l e s on a wind t u n n e l model. I n addi- t i o n t o t h e s t a t i c s near t h e e x i t , there were f o u r r a k e s w i t h t h r e e probes each l o c a t e d a t t h e s i m u l a t e d compressor face. These were designed t o measure p o s s i b l e flow d i s t o r t i o n caused by t h e propfan n a c e l l e s l i p - stream.

A n i n t e r n a l d r a g computation procedure f o r t h e Spey n a c e l l e s was developed and allows f o r i n l e t t o t a l p r e s s u r e s d i f f e r e n t from f r e e s t r e a m .

Entry t o t a l p r e s s u r e s were obtained by working forward from t h e r a k e s assuming t h e o n l y d u c t l o s s t o be t h a t of i n t e r n a l f r i c t i o n . A d u c t p r e s s u r e loss f a c t o r w a s i n c l u d e d f o r t h e e f f e c t of t h e r a k e s themselves.

The c o s i n e e f f e c t on t h e recovered e x i t t h r u s t produces a s l i g h t angle-of- a t t a c k (Alpha) and n a c e l l e i n c i d e n c e e f f e c t on i n t e r n a l d r a g and t h e r e is a l s o a l i f t e f f e c t which i s a r e l a t i v e l y small percentage of the t o t a l l i f t , b u t i s n e v e r t h e l e s s a s i g n i f i c a n t number. Both the i n t e r n a l d r a g and t h e i n t e r n a l l y g e n e r a t e d incremental l i f t increments have been com- puted.

INTERNAL DRAG, POWERED NACELLE The i n l e t f o r t h e powered n a c e l l e is followed by flow-through d u c t w i t h a f a i r l y complex i n t e r n a l shape. It h a s a 5-tube t o t a l p r e s s u r e rake assembly a t t h e i n l e t e n t r y t o measure t h e t o t a l pressure. S i n c e t h e i n l e t i s l o c a t e d behind t h e prop, t h e t o t a l temperature as w e l l as the t o t a l p r e s s u r e of t h e stream is a f f e c t e d . The i n t e r n a l d r a g computation f o r t h i s n a c e l l e i n c l u d e s t h i s temperature change. The procedure a l s o accounts f o r f r i c t i o n , expansion, and t u r n i n g l o s s , which o c c u r s i n s i d e the duct.

APPENDIX E SYMBOLS AF Balance a x i a l f o r c e

BL Butt-Line - lateral p o s i t i o n measured from a i r c r a f t c e n t e r l i n e

Drag c o e f f i c i e n t cD R o l l i n g moment c o e f f i c i e n t c1 L i f t c o e f f i c i e n t CL C or CM P i t c h i n g moment c o e f f i c i e n t m P i t c h i n g moment c o e f f i c i e n t a t zero l i f t Yawing moment c o e f f i c i e n t ‘ n C P r e s s u r e c o e f f i c i e n t P c * P r e s s u r e c o e f f i c i e n t corresponding t o Mach 1.0 P Power c o e f f i c i e n t cP Thrust c o e f f i c i e n t Ct C Side f o r c e c o e f f i c i e n t Y d Prop diameter D Drag FRL Fuselage r e f e r e n c e l i n e H T o t a l p r e s s u r e Freestream t o t a l pressure HO HP Horsepower J P r o p e l l e r advance r a t i o , V 0 b d L L i f t M Mach number Freestream Mach number MO MAC Mean aerodynamic chord x 1 SYMBOLS (CONTINUED) N F Balance normal f o r c e NT Nacelle I n c i d e n c e Angle PW Balance p i t c h i n g moment P S t a t i c p r e s s u r e Freestream s t a t i c p r e s s u r e PO 9 Dynamic p re s s u r e Freestream dynamic p r e s s u r e RM Balance r o l l i n g moment Rrl N e u t r a l p o i n t SF Balance s i d e f o r c e SQRT Square r o o t TC o r TC Thrust c o e f f i c i e n t U A x i a l v e l o c i t y i n p r o p e l l e r c o o r d i n a t e s V Lateral v e l o c i t y I n p r o p e l l e r c o o r d i n a t e s W Vertical v e l o c i t y i n p r o p e l l e r c o o r d i n a t e s Freestream v e l o c i t y vO R e s u l t a n t v e l o c i t y "R d i s t a n c e from l e a d i n g edge 1 Chordwise wing p o s i t i o n (

x/c

wing chord d i s t a n c e from n a c e l l e s t a t i o n s Lengthwise n a c e l l e p o s t t i o n ( X/L 1 nacelle length YM Balance yawing moment Greek Symbols a A i r c r a f t a n g l e of a t t a c k P r o p e l l e r a n g l e of a t t a c k C Y t

P A i r c r a f t yaw a n g l e

B P r o p e l l e r b l a d e p i t c h a n g l e SYMBOLS (CONTINUED) Aileron d e f l e c t i o n Rudder d e f l e c t i o n S t a b i l i z e r d e f l e c t i o n 7) Spanwise wing s t a t i o n i n p e r c e n t of semi-span dimension CL Azimuth a n g l e i n p r o p e l l e r p l a n e w P r o p e l l e r r o t a t i o n a l speed 4 * REFERENCES 1. H a c k e t t , J. E., C. G. P h i l l i p s , and D. E. L i l l e y , "Three-Dimensional Wake Flow Measurements f o r a Wing and a B l u f f , Car-Like Body," Lockheed-Georgia Company, LG81ER0201, Submitted t o t h e Georgia I n s t i t u t e of Technology, August 1981.

2. C A T I A APAR C o o r d i n a t o r , I B M Corporation, "Computer-Graphics Aided Three-Dimensional I n t e r a c t i v e A p p l i c a t i o n (CATIA) User Manual, Program Numbers 5796-PQG, 5796-PQH, 5796-PQJ, and 5796-PQL, Second E d i t i o n , " 8820-2629-4, March 1984.

3. Youngren, H. H., E. E. Bouchard, R. M. Coopersmith, and L . R. Miranda, f l Comparison of Panel Method Formulations and Its I n f l u e n c e on t h e Development of QUADPAN, an Advanced Low Order Method," AIAA-83-1827, A I A A Applied Aerodynamics Conference, 3-15 J u l y 1983.

4. A l j a b r i , A. S., and J. M. Swearengen, "Analysis of Mach Number 0.8 Turboprop S l i p s t r e a m Wing/Nacelle I n t e r a c t i o n s , " NASA CR166419 (prepared by Lockheed-Georgia Company and i s s u e d as LG82ER01691, December 1982.

5 . A l j a b r i , A. S., and B. H. L i t t l e , Jr., "High Speed Wind Tunnel Tests of t h e PTA A i r c r a f t , " SAE 861744, Aerospace Technology Conference and E x p o s i t i o n , 13-16 October 1986.

6. A l j a b r i , A. S., and C. A. Hughes, "Wind Tunnel I n v e s t i g a t i o n of Pro- p e l l e r S l i p s t r e a m I n t e r a c t i o n w i t h Nacelle/Wing/Flap Combinations," AGARD Paper 21, p r e s e n t e d t o AGARD Symposium on Aerodynamics and Acoustics of P r o p e l l e r s , Toronto, Canada, 1-4 October 1984.

PREXXDING PAGE NOT F T L M E E D

Report Documentation Page

1. Report No. 2. Government Accession No.

3. Recipient's Catalog No.

NASA CR-182121 4. Title and Subtitle 5. Report Date

PROPFAN TEST ASSESSMENT

May 1988

TESTBED AIRCRAFT STABILITY AND CONTROL/PERFORMANCE

6. Performing Organization Code

1/9-SCALE WIND TUNNIX TESTS

I 7. Authorts) 8. Performing Organization Report No.

B. H. Little, Jr., K . H. Tomlin, LG88ER0056 A . S . Aljabri, and C . A . Mason I O . Work Unit No.

9. Performing Organization Name and Address 11. Contract or Grant No.

Lockheed Aeronautical Systems Company NAS3-24339 8 6 South Cobb Drive Marietta, Georgia 30063 13. Type of Report and Period Covered Contractor Report 2. Sponsoring Agency Name and Address Task 7 Final National Aeronautics and Space Administration 14. Sponsoring Agency Code Lewis Research Center Cleveland, Ohio 44135-3191 I 5. Supplementary Notes

Project Manager, John B. Whitlow, Propfan Test Assessment Project

NASA Lewis Research Center 6 Abstract One-ninth scale wind tunnel model tests of the Propfan Test Assessment (PTA) aircraft were performed in three different NASA facilities. Wing and propfan nacelle static pressures, model forces and moments, and flow field at the propfan plane were measured in these tests. Tests started in June 1985 and were completed in January 1987.

These data were needed to assure PTA safety of flight, predict PTA performance, and validate analytical codes that will be used to predict flow fields in which the propfan will operate.

7. Key Words (Suggested by Author@)) 18. Distribution Statement Aerodynamics

Unclassified - Unlimited

Wind Tunnel Testing Subject Category 05 Propfan Aircraft Design 9. Security Classif. (of this report) 20 Security Classif. (of this page) 21. No of pages 22. Price' Unclassified Unclassified I I