Wind tunnel and ground static investigation of a large scale model of a lift/cruise fan V/STOL aircraft

Section

MDC A4318 TABLE OF COXTENTS Section Title .

ii SumfARY . . . . . . . . . . . . . . . . . . . . . . . . . . . .

W LIST OF FIGURES .

SYMBOLS AYD NOMENCLATURE . . . . . . . . . . . . . . . . . . . xv

IX'RODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . 1-1

1 .

1 . 1 Background . . . . . . . . . . . . . . . . . . . . . . . . 1-1

1 . 2 40' x 8 0 ' Wind Tunnel Tests . . . . . . . . . . . . . . . 1-2

1 . 3 Outside Static Tests . . . . . . . . . . . . . . . . . . . 1-2

2 . 2-1 HODEL DESCRIPTIO23 . . . . . . . . . . . . . . . . . . . . . . .

2-4 2 . 1 Fuselage . . . . . . . . . . . . . . . . . . . . . . . . . .

2-4 2 . 2 Wing . . . . . . . . . . . . . . . . . . . . . . . . . . .

2-4 2 . 3 Empennage . . . . . . . . . . . . . . . . . . . . . . . .

2- 7 2.4 Air Induction System . . . . . . . . . . . . . . . . . . .

2-11 2 . 5 Gas Generator anci Turbotip Fan System . . . . . . . . . .

2- ll 2 . 6 Thrust Vectoring System . . . . . . . . . . . . . . . . .

3-1 3 . HODEL INSTRUMENT,4TION . . . . . . . . . . . . . . . . . . . . .

3-1 3 . 1 Airframe Instrumentation . . . . . . . . . . . . . . . . .

3-1 3 . 2 Propulsion System Instrumentation . . . . . . . . . . . .

4-1 4 . WIND TUNNEL TEST FACILITY . . . . . . . . . . . . . . . . . . .

4-1 4 . 1 Test Arrangement . . . . . . . . . . . . . . . . . . . . .

4-9 4.2 Data Acquisition System . . . . . . . . . . . . . . . . .

OUTSIDE STATIC TEST FACILITY . . . . . . . . . . . . . . . . . . 5-1

5 .

5-1 5 . 1 Test Arrangement . . . . . . . . . . . . . . . . . . . . .

5 . 2 Data Acquisition System . . . . . . . . . . . . . . . . . 5-7

WIND TUNIEL TEST P R O G W Y . . . . . . . . . . . . . . . . . . . 6-1

6 .

6.1 Test Conditions . . . . . . . . . . . . . . . . . . . . . 6-1

6.2 Test Procedures . . . . . . . . . . . . . . . . . . . . . 6-4

6.3 Data Reduction . . . . . . . . . . . . . . . . . . . . . . 6-6

7-1 OUTSIDE STATIC TEST PROGRAM . . . . . . . . . . . . . . . . . .

7 .

7-1 7 . 1 Test Conditions . . . . . . . . . . . . . . . . . . . . .

7-3 7.2 Test Procedures . . . . . . . . . . . . . . . . . . . . .

7-6 7.3 Data Reduction . . . . . . . . . . . . . . . . . . . . .

8-1 8 . WIND TUNNEL TEST RESULTS . . . . . . . . . . . . . . . . . . .

8-1 8.1 Propulsion System Static Callbration . . . . . . . . . . .

8-21 8.2 Propulsion System Performance . . . . . . . . . . . . . .

8.3 Powered Lift Configuration - Induced Lift and Drag

8-36 Characteristics . . . . . . . . . . . . . . . . . . . .

8.4 Powered L i f t Configuration . Pitching Yompnt

8-117 Characteristics . . . . . . . . . . . . . . . . . . . .

8.5 Powered Lift Configuration . Lateral-Directional

Characteristics . . . . . . . . . . . . . . . . . . . . 8-126

MCQONNELb AIRCRAFT COMPANY iii M ) ( : A4318 TABLE OF CONTEXTS (Continued) T i t l e Page

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8.6 Aerodynamic L i f t Configuration - L o n g i t u d i n a l

C h a r a c t e r i s t i c s . . '. . . . . . . . . . . . . . . . . . 8-152

8.7 Aerodynamic L i f t Configuration - L a t e r a l - D i r e c t i o n a l

C h a r a c t e r i s t i c s . . . . . . . . . . . . . . . . . . . . 8-173

8.8 A i r I n d u c t i o n System Performance. . . . . . . . . . . . . 8-191

9. OUTSIDE STATIC TEST RESULTS. . . . . . . . . . . . . . . . . . 9-1

9 . 1 Propulsion System C a l i b r a t i o n s . . . . . . . . . . . . . . 9-1

9 . 2 Ground E f f e c t s on A i r c r a f t L i f t Loss. . . . . . . . . . . 9-13

9.3 Grc;md E f f e c t s on I n l e t Reingestion . . . . . . . . . . . 9-37

9.4 Flow V i s u a l i z a t i o n Tests. . . . . . . . . . . . . . . . . 9-51A

10. CONCLUSIONS. . . . . . . . . . . . . . . . . . . . . . . . . . 10-1

11. LIST OF REFEREXCES . . . . . . . . . . . . . . . . . . . . . . 11-1

APPENDIX h 40' x 80' WIXD TUNNEL TEST SCHEDULE AXD BALMCE DATA . . . . . A- 1

APPENDIX B OUTSIDE STATIC TEST RW S U I W S . . . . . . . . . . . . . . . . B- 1

L i s t of Pages T i t l e ii through x v i i i 1-1 artd 1-2 2-1 through 2-17 3-1 through 3-17 4-1 through G-9 5-1 through 5-10 6-1 through 6-9 7-1 through 7-7 8-1 through 8-206 9-1 through 9-51A, 9-52 through 9-64 10-1 and 10-2 11-1 A-1 through A-22 B-1 through B-S MC0ONNal.L AIRCRAFT COMPANY i v MDC A4318 LIST OF FIGURES Nusber T i t l e .

Large S c a l e Powered Nodel of a L f f t l C r u i s e Fan 2- 1 V/STOL A i r c r a f t Concept . . . . . . . . . . . . . . . . . . . . .

2-2

Large S c a l e Powered Nodel P r o p u l s i o n System . . . . . . . . . . . . 2-2

2-3

Large S c a l e Powered Node1 . . . . . . . . . . . . . . . . . . . . . 2- 3

2-5 Large S c a l e L i f t / C r u i s e Fan A i r c r a f t Model . . . . . . . . . . . . 2- 4 2-6

L i f t / C r u i s e Fan I n l e t Design Geometry . . . . . . . . . . . . . . . 2- 5 2-8

L i f t / C r u i s s Gas Generator I n l e t Design Geometry 2- 6

. . . . . . . . . . 2-9

Gas Generator I n l e t S h i e l d Geometry . . . . . . . . . . . . . . . . 2- 7

2-10

Nose Fan I n l e t Design Geometry . . . . . . . . . . . . . . . . . . 2-8

2-12

Forward Engine I n l e t Design Geometry . . . . . . . . . . . . . . . 2-9

2-13 Gas Generator and T u r b o t i p Fan Design Characteristics 2- 10 2-14 . . . . . . .

L i f t / C r u i s e Unit Vectoring System Geometry . . . . . . . . . . . . 2- 11 2-15

Nose L i f t Unit Vectoring System Geometry . . . . . . . . . . . . . 2-12

2-16 3-1 Location of Wing S u r f a c e S t a t i c P r e s s u r e P o r t s . . . . . . . . . . 3-2

3-2 Location of Forward Fuselage S t a t i c P r e s s u r e P o r t s . . . . . . . . 3 3

3-3 P r o p u l s i o n System I n s t r u m e n t a t i o n a a t i o n a l e . . . . . . . . . . . . 3-5 3-4 L i f t / C r u i s e Fan I n l e t I n s t r u m e n t a t i o n . . . . . . . . . . . . . . . 3-6 3-7 3 5 Xose Fan I n l e t I n s t r u n e n t a t i o n . . . . . . . . . . . . . . . . . .

3-6 3-8 L i f t / C r u i s e Engine I n l e t I n s t r u m e n t a t i o n . . . . . . . . . . . . .

3-7 3-9 Forward Engine I n l e t I n s t r u m e n t a t i o n . . . . . . . . . . . . . . .

3- 8 3-10 Fan Face f n s t r u m e n t a t i o n . . . . . . . . . . . . . . . . . . . . .

3-9 Engine Face I n s t r u m e n t a t i o n . . . . . . . . . . . . . . . . . . . . 3-12

3- 10 Fan and T i p Turbine E x i t I n s t r u m e n t a t i o n . . . . . . . . . . . . . 3-13 3-11 Fan I n l e t and E x i t I n s t r u n e n t a t i o n 1 n q : a l l a t i o n . . . . . . . . . . 3-14 3-12 Engine E x i t I n s t r u m e n t a t i o n . . . . . . . . . . . . . . . . . . . . 3-15 3-13 Nozzle E x i t I n s t r u n e n t a t i o n . . . . . . . . . . . . . . . . . . . . 3-17

4- 1 40 x 80 Wind Tunnel Test F a c i l i t y . . . . . . . . . . . . . . . . . 4-2

4- 2 Large S c a l e Pcwered Nodel I n s t a l l e d i n 40 f t x 80 f t

Wind Tunnel . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-3

4- 3 Large S c a l e Powered Xodel I n s t a l l e d i n 40 f t x 80 f t

Wind Tunnel . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-4

4- 4 Large S c a l e Powered Hodel I n s t a l l e d i n 40 f t x 80 f t

Wind Tunnel . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-5

4- 5 Large S c a l e Powered Xodel I n s t a l l e d i n 4 0 i t x 80 f t

Wind Tunnel . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-6

4- 6 Large S c a l e Powered ?!ode1 I n s t a l l e d in 40 f t x 80 f t

Wind Tunnel . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-7

4- 7 40 x 80 Control Room Layout . . . . . . . . . . . . . . . . . . . . 4-8

5- 1 Outside S t a t i c Test F a c i l i t y P l a n V i e w . . . . . . . . . . . . . . 5-2

5-2 Node1 Ground Height V a r i a t i o n s . . . . . . . . . . . . . . . . . . 5-3

5- 3 Large Scale Powered Xodel I n s t a l l e d i n Outside S t a t i c

Test F a c i l i t y . . . . . . . . . . . . . . . . . . . . . . . . . . 5-4

5-4 Large S c a l e Powered >lode1 I n s t a l l e d i n Outside S t a t i c

Test F a c i l i t y . . . . . . . . . . . . . . . . . . . . . . . . . . 5-5

5- 5 Large S c a l e Powered Node? I n s t a l l e d i n Outside S t a t i c 3-6 Test F a c i l i t y . . . . . . . . . . . . . . . . . . . . . . . . . .

Data A c q u i s i t i o n Systems Flow Diagram . . . . . . . . . . . . . . .

5- 6 5-8 O u - s i d e S t a t i c Test F a c i l i t y C o n t r o l Trailer Layout . . . . . . . .

5- 7 5-9 MCDONNELL AIRCRAFT COMPANY MDC A4318 LIST OF FIGURES (Continued) Number T i t l e Equations and Constants f o r Modeling Direct Propulsion 6- 1

System Components . . . . . . . . . . . . . . . . . . . . . . . . 6-9

8.1-1 Left Lift/Cruise Unit Static Thrust . . . . . . . . . . . . . . . . 8-3

8.1-2 Right Lift/Cruise Unit S t a t i c Thrust . . . . . . . . . . . . . . . 8-4

8.1-3 Nose L i f t Unit S t a t i c Thrust . . . . . . . . . . . . . . . . . . . 8-5

8-6 8.1-4 Nose L i f t Unit S t a t i c Thrust . . . . . . . . . . . . . . . . . . .

8-7 8.1-5 L e f t Lift/Cruise Unit I d e a l Thrust . . . . . . . . . . . . . . . .

8.1-6 Right L i f t l C r u i s e Unit I d e a l Thrust . . . . . . . . . . . . . . . . 8-8

8.1-7 Nose L i f t Unit I d e a l Thrust . . . . . . . . . . . . . . . . . . . . 8-9

8.1-8 Nose L i f t Unit I d e a l Thrust . . . . . . . . . . . . . . . . . . . . 8-10

8.1-9 Thrust Calibration Ccefficient Determination . . . . . . . . . . . 8-11

8.1-10 L i f t l C r u i s e Unit Thrust Coefficients . . . . . . . . . . . . . . . 8-12

8-13 8.1-11 Nose L i f t Unit Thrust Coefficients . . . . . . . . . . . . . . . .

8.1-12 ',eft Lift/Cruise Unit Thrust Vector Angles . . . . . . . . . . . . 8-14

8-15 8.1-13 Right Lift/Cruise Unit Thrust Vector Angles . . . . . . . . . . . .

8-16 8.1-14 Nose L i f t Unit Thrust Vector Angles . . . . . . . . . . . . . . . .

8-17 8.1-15 Lift/Cruise Unit Thrust Vector Angles . . . . . . . . . . . . . . .

Nose L i f t Unit Thrust Vector Angles . . . . . . . . . . . . . . . . 8-18

8.1-16 S t a t i c Pitching Moment Variation with Lift/Cruise Unit 8.1-17 8-19 . . . . . . . . . . . . . . . . . . .

Geometric Deflection Angle 8.1-18 S t a t i c Pitching Moment Variation with Nose L i f t Unit 8-20 Geometric Deflection Angle . . . . . . . . . . . . . . . . . . .

8-22 8.2-1 Lift/Cruise Unit Typical Performance C h a r a c t e r i s t i c s . . . . . . .

8-23 8.2-2 Nose L i f t Unit Typical Performance C h a r a c t e r i s t i c s . . . . . . . .

8-24 8.2-3 Angle of Attack Effect on Lift/Cruise Unit Performance . . . . . .

8-25 8.2-4 Angle of Attack Effect on Nose L i f t Unit Performance . . . . . . .

8.2-5 Effect of High Angles of Attack on Propulsion System 8-26 Performance . . . . . . . . . . . . . . . . . . . . . . . . . . .

Forward Speed Effect on LiftICruise Unit Performance . . . . . . . 8-27

a . 2-6

Forward Speed Effects on Various Nose Unit Vector Angles . . . . . 8-28

8.2-7 8.2-8 Comparison of Forward Speed Effects on the Lift/Cruise vs .

8-29 Nose L i f t Units . . . . . . . . . . . . . . . . . . . . . . . . .

8-30 8.2-9 Xass Flow Variations w i t h Forward Speed . . . . . . . . . . . . . .

Mass Flow Variations with Forward Speed . . . . . . . . . . . . . . 8-31

8.2-10 8-32 8.2-11 Mass Flow Variations with Forward Speed . . . . . . . . . . . . . .

8-33 8.2-12 Effect of Forward Speed on J e t Velocity Ratio . . . . . . . . . . .

8-34 8.2-13 Effect of Forward Speed on J e t Velocity Ratio . . . . . . . . . . .

8-35 8.2-14 Effect of Forward Speed on J e t Velocity Ratio . . . . . . . . . . .

8.3-1 L i f t vs Dynamic Pressure -r = 0'. s = 0'. BLC = 23'.

5 m = 43'. SJ 2 0 . 1 ' . . . . . . . . . . . . . . . . . . . . . . 8-38

8.3-2 Drag vs Dynamic Pressure 6~ = 0 . . cr = 0'. 6 ~ c = 23'.

6 ~ t 0 43'. 8 j = 2 0 . 1 ' . . . . . . . . . . . . . . . . . . . . . . 8-39

8.3-3 L i f t vs Dynamic Pressure 6~ = 0'. a = 8' ~ L C = 23..

5 m 43'. 85 ID 2 0 . 1 ' . . . . . . . . . . . . . . . . . . . . . . 8-40

Drag vs Dynamic Pressure SH = 0'. x = 8.. BLC 23..

8.3-4

6NL 43'. a j 2 0 . 1 ' . . . . 8-41

MCQONNBLL AJRCRAPT COMPANY vi MDC A4318 LIST OF FIGURES (Continued) Number T i t l e Page

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8.3-30 Drag vs Dynamic Pressure 6g = O', a = 16", 6~ = go',

6 ~ ~ 9 0 ' . 8 ~ ~ 8 4 . 7 ' . . . 6 . . . . 8-67

8.3-31 Induced L i f t and Drag Parameters vs Jet Velocity Ratio

6~ a O', 6 ~ c 23', 6 a 43', 8j = 20.1' . . . . . . . . . . . . 8-68

8.3-32 Induced L i f t and Drag Parameters vs Jet Velocity Ratio

8H o', 6 ~ c m 38", 6m 43', 8J 29.2O . . . . . . . . . . . . 8-69

8.3-33 Induced L i f t and Drag Parameters vs Jet Velocity Ratio

6g O', 6 ~ c = 56', 6 m 43', 85 44.5" . . . . . . . . . . . . 8-70

8.3-34 Induced L i f t and Drag Parameters vs Jet Velocity Ratio . . . . . . . . . . . .

6~ O', 6Lc = 71', 6 m = 55', BJ 59.8' 8-71 8.3-35 Induced L i f t and D r a g Parameters vs Jet Velocity Ratio 90', B j = 84.7' . . . . . . . . . . . .

8-72 6H = O", 6LC = goo, 8.3-36 Summary of Propulsion uced Effects at 0' Angle of Attack 8-73 Three Fan Operation, Horizontal T a i l On . . . . . . . . . . . . .

8.3-37 L i f t vs D y n d c Pressure, Horizontal T a i l Off

a 0'9 6Lc 23'9 6 m 43', 8J 20.1' . . 8-74

8.3-38 Drag vs Dynamic Pressure, Horizontal T a i l Off

a O', 6Lc a 23', 6 m = 43', BJ = 20.1' . . . . . . . . . . . . 8-75

8.3-39 L i f t vs Dynamic Pressure, Horizontal T a i l Off

a 8'. 6 ~ c 23', 6 m 43', BJ p 20.1' . . . . . . . . . . . . 8-76

8.3-40 Drag vs D y n d c Pressure, Eorizontal Tail Off

u 8' d ~ c 0 23', 6m 43', 85 20.1' . . 8-77

8.3-41 L i f t vs Dyncrmic. Pressure, Horizontal Tail Off

23', 6m 43O, BJ = 20.1' . . . . . . . . . . . . 8-78

8.3-42 Drag w Dyn c Pressure, Horizontal T a i l Off

a = 1 6 0 9 2

0 23", 6m p 43', BJ * 20.1' . 8-79

a 0 16', 6Lc 8.3-43 L i f t vs Dynamic Pressure, Horizontal T a i l Off

= 00, 6Lc = 560, 6NL = 430, eJ = 44.50 . . . . . . . . . . . . 8-80

8.3-44 Drag vs Dynamic Pressure, Horizontal T a i l Off

. . . 8-81

56', 6 m ~ 43', 8j = 44.5'

01 0"s 6Lc 8.3-45 L i f t vs Dynamic Pressure, Horizontal Tail Off

a = 8', 6 ~ c . 56", 6m = 43', 85 = 44.5' . . . . . . . . . . . . 8-82

8.3-46 Drag vs Dynamic Pressure, Horizontal Tail Off

a 8', 6 ~ c = 56', 6m 43', 8J 44.5' . . . . . . . . . . . . 8-83

8.3-4' L i f t vs Dynamic Pressure, Horizontal Tail Off

a 16', 6Lc 56', 6m 43', B j 0 44.5' . . 8-84

8.3-48 Drag vs Dynamic Pressure, Horizontal Tail Off

a 16", 63

8-85 56', 6 m * 43', B j 0 44.5' .

8.3-49 L i f t vs Dyna c Pressure, Horizontal T a i l Off

a = O ' , 6 ~ c = go', 6m . go', BJ = 84.7' . . . . . . . . . . . . 8-86

8.3-50 Drag vs Dynamic Pressure, Horizontal Tail Off

a O ' , 6 ~ c . 90', 5 m 0 90°, OJ 84.7' . . . . . . . . . . . . 8-87

8.3-51 L i f t vs Dynamic Pressure, Horizontal Tail Off

a = 8', 6Lc 90', 6 m = go', BJ . 84.7' . . . . . . . . . . . . 8-88

8.3-52 Drag vs Dynamic Pressure, Horizontal T a i l Off

a 8 O , 6 9 c = 90", bm = 9 0 ° , OJ = 84.7' . . . . . . . . . . . . 8-89

8.3-53 L i f t vs Dynaaic Pressure, Horizontal Tail Off a

le', 6Lc . go', 6m . go', BJ = 84.7' . . . . . . . . . . . . 8-9Q

8.3-54 Drag vs Dynamic Pressure, Horizontal Tail Off

CZ . 16', 6Lc . go', S m . go', OJ . 84.7' . . . . . . . . . . . . 8-91

MGDONNELL AtReff4 PT COMPANY v i i i MDC A4318 LIST O F FIGURES (Continued) T i t l e Page

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8.3-55 Induced L i f t and Drag Parameters vs Jet Velocity Ratio Horizontal T a i l Off 6 L ~ = 23', 6m = 43', 8J = 20.1'

. . . . . . 8-92

8.3-56 Induced L i f t and Drag Parameters vs Jet Velocity Ratio Horizontal T a i l Off 6Lc = 56', 6m = 43', 8J = 44.5' . . . . . .

8-93 8.3-57 Induced L i f t and Drag Parameters vs Jet Velocity Ratio . . . . . .

Horizontal T a l l Off 6 ~ c = 90°, 6 m * go", 95 = 84.7' 8-94 8.3-58 Sunrmary of Propulsion Induced Effects a t 0' Angle of Attack

Three Fan Operation, Horizontal T a i l Off . . . . . . . . . . . . 8-95

8.3-59 L i f t vs Dynamic Pressure, Flow Survey Rake On

2 = O', dLC = O', 6NL = 0' I n l e t s Covered, OJ = 1" . . . . . . . 8-96

8.3-60 'irag vs Dynamic Pressure, Flow Survey Rake O n 8-9 7 a = O', 6 ~ c = O', 6 m = 0' I n l e t s Covered, BJ = 1" . . . . . . .

8.3-61 L i f t vs Dynamic Pressure, Flow Survey Rake On

. . . . . . 8-98

a - O', dLC = 56", 6NL = 0' I n l e t s Covered, 8J = 47"

a. 3-62

Drag vs Dynamic Pressure, Flow Survey Rake On

a = O', 6Lc = 5 6 ' , 6m = 0' I n l e t s Covered, 8J = 47' . . . . . . 8-99

8.3-63 L i f t vs Dynamic Pressure, Flow Survey Rake On

a = O', 6 ~ c = go', 6m = 0' I n l e t s Covered, BJ = 84" . . . . . . 8-100

8.3-64 Drag vs Dynamic Pressa?, €low Survey Rake On

= 90°, & = 0 ' I n l e t s Covered, 8 = 84' . . . . . . 8-101

8.3-65 Summary a 0 ° , of 6kC ropulsion Educed Effects a t 0' Angie of Attack

8-102 Two Fan Operation, Flow Survey Rake O n . . . . . . . . . . . . .

8.3-66 Summary of Propulsion Induced Effects at 0' Angle of Attack 8-103 Two Fan Cperation, Flow Survey Rake O n . . . . . . . . . . . . .

8.3-67 Effect of Nose L i f t Unit on Total Aerodynamic L i f t

6 ~ ~ = 5 6 ' , a = O o . . . . . . . . . . . . . . . . . . . . . . . . 8- 104

8.3-68 Effect of Nose L i f t Unit on Total Aerodynamic Dra;?

dLC = 56', a = 0' . . . . . . . . . . . . . . . . . . . . . . . . 8-105

8.3-69 Effect of Nose L i f t Unit on Total Aerodynamic L i f t

6LC'90°, a = O 0 . . . . . . . . . . . . . . . . . . . . . . . . 8-106

8.3-70 Effect of Nose L i f t Unit on Total Aerodynamic Drag

6LC'90°, a = O 0 , . . . . . . . . . . . . . . . . . . . . . . . 8- 10 7

8.3-71 L i f t vs Dynamic Pressure, Flow Survey Rike On

a O', 6 ~ c = O o , 6m = S O ' , BJ m 14.8' . . . . . . . . . . . . . 8- 108

8.3-72 Drag vs Dynamic Pressure, Flow Survey Rake O n

a - O', 6LC = o', 6m Soo, 8j 14.6' . . 8- 109

8.3-73 L i f t vs Dynamic Pressure, Flow Survey Rake On

a = o', 6 ~ c = o', 6~ = 70°, 8 j = 20.3' . . . . . . . . . . . . . 8-110

8.3-74 Drag vs Dynamic Pressure, Flow Survey Rake O n a = O ' , 6 ~ c = O', 6~ = 70', BJ 2 0 . 3 ' . . . . . . . . . . . . .

8-111 8.3-75 L i f t vs Dynamic Pressure, Flow Survey Rake O n 8- 112 a = O', 6 ~ c = O o , 6m = 9 0 ° , BJ 23.8' . . . . . . . . . . . . .

8.3-76 Pressure, Flow Survey Rake On Drag vs Dynamic

Q O', 6 ~ c O ' , 5 m = 9 0 ° , 8j 23.8' . . . . . . . . . . . . . 8-113

8.3-77 Effect of Nose L i f t Unit cn Total Aerodynamic L i f t

6 ~ ~ = 0 ~ , a = O ' . . . . . . . . . . . . . . . . . . . . . . . . 8-114

8.3-78 Effect of Nose L i f t Unit on Total Aerodynamic L i f t

6 L C = O 0 , a = o 0 . . . . . . . . . . . . . . . . . . . . . . . . 8- 115

8.3-79 Summary of Propulsion Induced Effects a t 0' Angle of Attack

Effect of Nose Unit, Horizontal T a i l Off . . . . . . . . . . . . 8-116

MCDONNmLL AIUCUAPT COMPANY ix MDC A4318 LIST OF FIGURES (Continued) T i t l e Page Number

-

Pitching Moment vs Nose L i f t Unit Geometric Deflection, 8.4-1

Flow Survey Rake On 6 ~ c = O " , a = 0" . . . . . . . . . . . . . . 8-118

Pitching Moment vs Nose L i f t Unit Geometric Deflection, 8.4-2

Horizontal T a i l O f f a = O " , 6 ~ c = 56' . . . . . . . . . . . . . . 8-119

8.4-3 Pitching Moment vs Nose L i f t Unit Geometric Deflection,

Horizontal T a i l Off a = O", 6 ~ c = 90' . . . . . . . . . . . . . . 8-120

8.4-4 Pitching Moment vs Dynamic Pressure

a = 0", 6Lc 0 23", 6m = 43', 05 = 20.1' . . . . . . . . . . . . 8-121

8.4-5 Pitching Moment vs Dynamic Pressure

a = O", 6 ~ c 0 38", 6m = 43", 05 5 29.2' . . . . . . . . . . . . 8-122

8.4-6 Pitching Moment vs Dynamic Pressure

a = O", 6 ~ c = 56", 6 m - 43", 85 = 44.5' . . . . . . . . . . . . 8-123

8.4-7 Pitching Moment vs Dynamic Pressure

59.8' . . 8-124

a = O", dLC 71', 6 m 5 5 " , 0 j 8.4-8 Pitching Moment vs Dynamic Pressure

. . . . . . . . . . . . 8-125

a = O", 6Lc = 90°, 6 m = 90°, 05 - 84.7'

8.5-1 Yawing Moment vs Angle of S i d e s l i p

6~ - O", a = O', 6 ~ c = 23", 6 m = 43", BJ 0 20.1", q 5 12.4 PSF . 8-127

8.5-2 Rolling Moment and Side Force vs Angle of S i d e s l i p

6~ = O " , a - O', 6 ~ c = 23", 6 m = 43", 05 = 20.1", q = 12.4 PSF . 8-128

8.5-3 Yawing Moment vs Angle of S i d e s l i p

6~ = o", a 5 o", 6 ~ c 23", 6 m 0 43", 8J = 20.1", q 5 19.4 PSF 8-129

8.5-4 Rolling Moment and Side Force vs Angle of S i d e s l i p 8-130 6~ = o", a = o", 6 ~ c = 23", 6 m = 43'1 8J 5 20.1", q 19.4 PSF .

8.5-5 Yawing Moment vs Angle of S i d e s l i p

6~ O', 0 O " , 6 ~ c 56", 6 m 0 43", 05 44.5", q 0 3.2 PSF . 8-131

8.5-6 Rolling Moment and Side Force vs Angle of S i d e s l i p 3.2 PSF 8-i32

6~ = O " , CL O " , 6 ~ c 0 56', 6 m 43", 05 = 44.5", q

8.5-7 Yawing Moment vs Angle of S i d e s l i p

6~ o', a o", 6 ~ c 56"s 6 m a 43'9 8J = 44.5'9 q = 7.1 PSF 8-133

8.5-8 Rolling Moment and Side Force vs Angle of S i d e s l i p

o", 6 ~ c = 56", 6m m 43", 85 0 44.5", q 0 7.1 PSF . 8-134

6~ = O', a 8.5-9 Yawing Moment vs Angle of S i d e s l i p

6~ - O ' , a = O " , 6 ~ c 56", 6 m 0 43", 05 0 44.5', q = 12.2 PSF . 8-135

8.5-10 Rolling Moment and Side Force vs Angle of S i d e s l i p

6 12.2 PSF . 43", 85 0 44.5", q = O', a = o o , 8-136

8.5-11 Yawfng Moment vs Ang e of Sides1 p 6€c 56"9 6f

. 8-137

G H O", a = 0", 6 ~ c = 90', 6~ = go', 05 a 84.7", q = 1.4 PSF 8.5-12 Roll.ing Moment and Side Force vs Angle of S i d e s l i p

SH 0 O', a = O " , 6 ~ c 0 90", 6m = 9 0 ° , BJ 0 84.7', q = 1 . 4 PSF . 8-138

8.5-13 Yawing Moment vs Angle of S i d e s l i p

3 8 O', a O " , 6 ~ c 0 9 0 ° , ~ N L = 90°, 85 84.7', q 0 3.3 PSF . 8-139

a. 5-11 Rolling Moaent and Side Force vs Angle of S i d e s l i p

5H 0 o", c1 = o", 6Lc 0 go', go', 8J 0 84.7'9 q a 3 . 3 PSF 8- 140

5 m

8.5-15 Yawing Moment vs Angle of S i d e s l i p

6H m 0'9 U o", ~ L C go", ~ N L L: ' O o , 8J 84.7'9 4 7.2 PSF 8-141

8.5-16 Rolling Moamt and Side Force vs Any: of S i d e s l i p

6~ o', 1 0 o", 6 ~ c go', 6m 0 r J " , 6J 0 84.7", 4 m 7 . 2 PSF . 8-142

MCDONNKLL AJRCHAPT COMPANY X MDC A4318 LIST OF FIGURES (Contir.*ied) Number T i t l e Page

-

8.5-17 Three Fan C o n f i g u r a t i o n D i r e c t i o n a l C h a r a c t e r i s t i c s 6 ~ = 0 " , a = O . . . . . . . . . . . . . . . . . . . . . . . . . .

8-143 8.5-18 Three Fan C o n f i g u r a t i o n D i r e c t i o n a l C h a r a c t e r i s t i c s

6 H 0 0 " , O E 0 " . . . . . . . . . . . . . . . . . . . . . . . . . 6-144

8.5-19 Three Fan C o n f i g u r a t i o n Lateral C h a r a c t e r i s t i c s

6 H P 0 " , Q 1 0 " . . . . . . . . . . . . . . . . . . . . . . . . . 5-3 45

8.5-20 Three Fan C o n f i g u r a t i o n Lateral C h a r a c t e r i s t i c s

6 ~ - O o , a = 0 . . . . . . . . . . . . . . . . . . . . . . . . . . f -146

8.5-21 Three Fan C o n f i g u r a t i o n S i d e Force Characteristics

6 H m o 0 , a r n o 0 . . . . . . . . . . . . . . . . . . . . . . . . . 8-147

8.5-22 Three Fan C o n f i g u r a t i o n S i d e Force C h a r a c t e r i s t i c s 6 ~ p ~ o , ~ p 8-148 8.5-23 E f f e c t of A i l e r o n Deflection on L a t e r a l - D i r e c t i o n a l C h a r a c t e r i s t i c s , q = 7.1 PSF

6~ = O " , a - O", 6 ~ c = 5 6 " , 6m 43', BJ = 44.5' . . . . . . . . 8-149

8.5-24 E f f e c t of A i l e r o n D e f l e c t i o n on L a t e r a l - D i r e c t i o n a l C h a r a c t e r i s t i c s , q = 12.3 PSF 6 H o", o", 6LC 56', 6m 43'9 8J 44.5" .

8- 15 0 8.5-25 E f f e c t of Rudder D e f l e c t i o n on L a t e r a l - D i r e c t i o n a l C h a r a c t e r i s t i c s

6~ 0 O " , 6Lc 3 8 " , 6 n 4 3 " , O j 29.2' . . . . . . . . . . . . 8- 15 1

8.6-1 L i f t vs Angle of A t t a c k

6~ 0 0 " , q 0 34.2 PSF, 6Lc 0 O " , 6m SEALED, 03 1" . 8- 15 5

8.6-2 Drag vs Angle of Attack

6~ = O " , q 34.2 PSF, 6 ~ c O ' , 6 m * SEALED, BJ = 1" . . . . . 8- 15 6

8.6-3 P i t c h i n ? Moment v s Angle of Attack

6~ 0 O ' , q 34.2 PSF, 6 ~ c 0 O " , 6 m SEALED, O j 1 ' . . . . . 8-15 7

8.6-4 L i f t vs Angle of Attack, H o r i z o n t a l T a i l Off

1" . . . . . . . . . . 8-158

q = 34.2 PSF, 6 ~ c = O " , 6 m = SEALED, 8j

8.6-5 Drag vs Angle of Attack, H o r i z o n t a l Tail Off

q = 34.2 PSF, 6 ~ c = 0". 6 m = SEALED, 0J 0 1 ' . . . . . . . . . . 8-159

8.6-6 P i t c h i n g Moment v s Angle of Attack, H o r i z o n t a l Tail Off q 0 34.2 PSF, 6 ~ c = O', 6m SEALED, 8J 1" 8- 160 8.6-7 L i f t vs Angle of Attack, Flow Survey Rake On

q = 34.2 PSF, 6 1" . . . = O ' , 6NL = 0" I n l e t s Covered, BJ 8-161

8.6-8 Drag vs Angle of kgtack, Flow Survey Rake On

q = 34.2 PSF, 6Lc = O ' , 6m - 0" I n l e t s Covered, 8J = 1" . . . . 8- 162

8.6-9 P i t c h i n g Moment vs Angle of Attack, F l a w Survey Rake On

q - 34.2 PSF, bLC = O ' , 6NL = 0 ' I n l e t s Covered, 8J 0 1" . . . . 8-163

8.6-10 E f f e c t of H o r i z o n t a l Tail on L i f t C o e f f i c i e n t vs Angle of A t t a c k q = 34.2 PSF, StC - 0", jNL = SEALZD, 9 j = 1 ' . . . . . . . . . .

8-164 8.6-11 E f f e c t of H o r i z o n t a l T a i l on L i f t C o e f f i c i e n t vs Drag C o e f f i c i e n t q = 34.2 PSF, 6 ~ c 8-165 O " , 6 m = SEALED, BJ * 1" . . . . . . . . . .

8.6-12 E f f e c t of H o r i z o n t a l Tail on P i L c t i n g Moment C o e f f i c i e n t vs Angle of Attack 5~ = O ' , q - 34.2 PSF, 6 ~ c = O ' ,

~ N L ~ S E A L E D , 8 j = l 0 . . . . . . . . . . . . . . . . . . . . . . 8-166

8.6-13 L i f t C o e f f i c i e n t v s Angle of Attack, Flow Survey Rake On

. . . . 0" I n l e t s Covered, 95 = lo 9 0 34.2 PSF, 6 ~ c 0 O " , 8-167

8,6-14

L i f t C o e f f i c i e n t v s Drag Coe f i c i c n t , Flow Survey Rake On 69

. . .

q - 34.2 PSF, 6tC 0 O o , €in - 0" I n l e t s Covered, a J = 1" 8- 16 8

MeDONNELL AIRCIRAtT COMPANV x i MDC A4318 LIST OF FIGURES (Continued) Number T i t l e Page

-

8.6-15 P-tching Moment C o e f f i c i e n t vs Angle of Attack, Flow Survey Rake On

. . . . q = 34.2 PSF, SLC = O " , 6 a = 0" I n l e t s Covered, BJ = 1" 8-169

8.6-16 E f f e c t of H o r i z o n t a l T a i l on L i f t C o e f f i c i e n t vs Angle of Attack

34.2 PSF, 6Lc O " , 6 m a SEALED, 05 a 1" . 8-170

q 8.6-17 E f f e c t of H o r i z o n t a l T a i l on L i f t C o e f f i c i e n t v* Drag C o e f f i c i e n t q 5 34.2 PSF, 6 ~ c O " , 6 m a SEALED, BJ L . . . . . . . . . .

8-171 8 6-18 E f f e c t of H o r i z o n t a l T a i l on P i t c h i n g Moment C o e f f i c i e n t v8 Angle

of Attack q 34.2 PSF, 6 ~ c = O " , 6m = SEALED, BJ = 1" . . . . 0-172

8.7-1 Side Force vs Angle of S i d e s l i p a = O", 6~ = O " , q 34.2 PSF, 6 ~ c = O " , Gm=SEALI?D, d j 1" . .

8-175 8.7-2 vs Angle of S i d e s l i p Yawing Moment

. 34.2 PSF, 6 ~ c = o", 6m * SELED, 8J a 1" 8-176

u = O " , 6~ O " , q 8.7-3 R o l l i n g Moment VB Angle of S i d e s l i p

u 0 O", 6~ 0 O", q SEALED, BJ a 1" . 34.2 PSF, 6 ~ c = O", 6m 8- 17 7

8.7-4 Side Force vs Angle of S i d e s l i p

a = 8 " , 6~ = O " , q = 34.2 PSF, 6 ~ c - O " , 6 m = SEALED, 6 j = 1" . 8-178

8.7-5 Yawing Moment v8 Angle of S i d e s l i p

U 8 " , 6~ = O " , q = 34.2 PSF, 6 ~ c = O " , 6m = SEALED, 8J 1" 8-179

8.7-6 R o l l i n g Moment vs Angle of S i d e s l i p Y - 8 " , 6~ = 0". q 34.2 PSF, 6 ~ c 0 O", 6 a SEALED, 8J 1" .

8-180 8.7-7 Side Force vs Angle o f S i d e s l i p u 0 1 6 " , 6 8 34.2 PSF, 6Lc * C " , 6 n 0 SEALED, 05 0 1" 8-181 O " , q 8.7-8 Yawing Moment vs Angle of S i d e s l i p

u 16", 6~ O " , q a 34.2 PSF, 6 ~ c O " , 6m 0 SEALZD, 8J 1" . 8-182

8.7-9 R o l l i n g Moment vs Angle of Sic'uslip

2 1 6 " , 6~ = O " , q 34.2 S" ', 6 ~ c O " , 6m SEALED, 8 j 0 1" . 8-183

8.7-10 E f f e c t of Angle of A t t a c k on LLde Force C o e f f i c i e n t vs Angle of Sides l i p

6~ O " , q e 34.2 PSF, 6 ~ c O " , 6m = SEALED, BJ = 1" . . 8-184

8.7-11 E f f e c t of Angle of A t t a c k on Yawing Moment C o e f f i c i e n t vs Angle of S i d e s l i p

6~ O " , q 34.2 PSF, 6 ~ c = O", 6 m 0 SEALED, BJ 0 1" 8-185

8.7-12 E f f e c t of Angle of Attack on R o l l i n g Moment C o e f f i c i e n t vs Angle of S i d e s l i p

6~ 0 O " , q 34.2 PSF, 6 ~ c 0 " , 6m 0 SEALED, 8 ; 1" . . . . 8-186

8.7-13 E f f e c t of A i l e r o n D e f l e c t i o n on L a t e r a l - D i r e c t i o n a l C h a r a c ~ e r i s t i c s u O " , 6H O " , q 34.2 PSF, 6Lc O " , 5m SEALED, BJ 1" .

8-187 8.7-14 E f f e c t of Aileron D e f l e c t i o n on L a t e r a l - D i r e c t i o n a l C h a r a c t e r i s t i c s

u O " , 6~ p 3". q 34.2 PSF, 6 ~ c a O", 6m SEALED, ?J = 1" 8-188

8.7-15 E f f e c t of Rudr'er D e f l e c t i o n on Side Force

6~ O", q p 34.2 PSF, 6Lc O", 5 m SEALED, 8, r. l o . . . . . 8-189

8 . ;-LU E f f e c t of Rudder D e f l e c t i o n on Yawing and R o l l i n g Mosents 6~ O " , q 34.2 PSF, 6Lc = O o , 6m SEALED, 05 ' 1 1 ' .

8-190 8.8-1 E f f e c t s of Angle of A t . a c k and Fan Speed on L i f t I C r u i s e Fan

I n l e t Performance . . . . . . . . . . . . . . . . . . . . . . . . 8- 19 4

8.8-2 E f f e c t s of Angle of 8 - 1 . . and S i d e s l i p on L i f L / C r u i s e Fan

I n l e t Performance . . . . . . . . . . . . . . . . . . . . . . . 8-195

8.8-3 E f f e c t s of S i d e s l i p Angle and Angle of Attack on L i f t / C r u i s e Fan and Gas Generator "t st Performance . . . . . . . . . . . . . . . .

8-196 -0ONNEl.I AIRCRAFT COMPANY x i i LIST OF FIGURES (Caatinued) T i t l e Imhr .

8.8-4 E f f e c t s of Foxward Speed and Angle of Attack on L i f t / C r u i b c

F m I a l e t P e r f o r m m c e . . . . . . . . . . . . . . . . . . 6-19 7

. . . .

8.8-5 Lift/Cruise Fan W e t Perforpedmce Suuumry . . . . . . . . . 8-198

. . . .

8.8-6 Effects of Augle of A t t a c k and E n g i n e Speed on L i f t / C r u i s e

Gas Generator falet Performance . . . . . . . . . . . . . . . . . 8-199

a . 8-1 Effects of Augle of Attack and S i d e s l i p on L i f t / C m i s e

Cas Generator Inlet Psrforsutnce . . . . . . . . . . . . . . . . . 8-200

8.8-8 Effects of &gle of Attack d Faa Speed on N o s e Fan I n l e t

P e r f o n m u c e g ...................... . . . . 8-201

8.8-9 Effects of Angle of Attack and Fan Speed on Nose! Fan Idlet

Perfowancg ....................... . . . . 8-202

8.8-10 Effects of Angle of Attack aad Faa Speed on Nose Fan I n l e t

Performance ....................... . . . . 8-203

8.8-U Nose Lift Unit Inlet Velocity Ratios . . . . . . . . . . . . . . . 8-204

8 . 8 - l 2 Effects of Sideslip Angle on Xose Fsm Inlet P e r f o m c e . . . . . . 8-205

a . 8-13 N o s e Fan I n l e t Perfowance Swnary . . . . . . . . . . . . . . . . 8-206

9-1 L e f t L i f t / C r u f s e U n i t Calibraticw Results . . . . . . . . . . . . . 5-3

e 2 Bight Lffc/Cruise Dait C a b r a t i o n Results . . . . . . . . . . . . 9-4

9-3 Rear Quarter View Showing L i f t / C r u i s e Vectoring Nozzles . . . . . . 9-5

9-4 Sose Lift U n i t Callbration Results . . . . . . . . . . . . . . . . 9-6

9- 5 . L i f t / C r u i s e Unit Thrust Coefficient Comparison . . . . . . . . . . 9-7

9-6 Nose L i f t Unit Thrust Coefff-cient CotPpar%soa . . . . . . . . . . . 9-8

9-7 L i f t / C r u i s e Untt Thrust Vector Angle Comparison . . . . . . . . . . 9-9

9-8 N o s e L i f t U n l t Thrust Vector Angle Comparison . . . . . . . . . . . 9-10

9-9 X376B FanHappingResults . . . . . . . . . . . . . . . . . . . . . 9-11

9-10 X376B F m Mapping W u l t s . . . . 7-12

9-11 EfCoct of Ground Height on Total Measured L i f t . . . . . . . . . . 9-14

9-12

Indiwidual U n i t I d e a l Thrust Measureatents . . . . . . . . . . . . . 9-15

9-13 9-16 Individual Unit I d e a l Thrust Measurements . . . . . . . . . . . . .

9-14 9-17 Individual Unit Ideal Thrust Measurements . . . . . . . . . . . . .

9-U . . . . 9-18

Effect of Ground Height on the I n d i d d u a l Unit Thrust . . .

9-16 . . . . . . . . . 9-20

Effect of Ground Height on Total L i f t and Thrust

9-17 Effect of I n l e t Shielding on T o t a l Measured L i f t . . . . . . . . . 9-21

9- I8 Effect of I n l e t Shielding on T o t a l k a s u r e d L i f t . . . . . . . . . 9-22

9-19 Effect of I n l e t Shielding on Total Measured L i f t . . . . . . . . . 9-23

9-20 Effect of Shieli Deflection Angle on Total Measured L i f t . . . . . 9-24

1-26 9-21 Effect of Grouud Height 011 Individual Unit Measured L i f t . . . . .

9-22 Effect of Ground Keight on Individual 3 n i t Measured L i f t . . . . . 9-27

9-23 Effect of Ground Height on Individual Unit Measured L i f t . . . . . 9-2C

9-24 Effect of Ground Height on Individual Unit I d e a l Thrust . . . . . . 9-29

9-25 Effect of Gzotmd Height on Individual Unit I d e a l Thrust . . . . . . 9-30

9-26 Effect of Ground Hefght on Individual Unit I d e a l Thrust . . . . . . 9-31

9-27 Effect of Ground Height on Individual U n i t L i f t and Thrust 9-32 . . . .

9-28 Effect of Ground Height on Individual Unit L i f t and Thrust 9-33 . . . .

9-29 9-34 Effect of Ground Height on Individual U d t L i f t and Thrust . . . .

9- 30

L i f t Loss I n Ground Effect Comparisons . . . . . . . . . . . . . . 9-36

9-31

Powered Model Temperature Instrumented Inlets . . . . . . . . . . . 9-38

9-32

Effect of Ground Height on I n l e t Reingestion . . . . . . . . . . . 9-39

9-33

Effect of Louver Deflection Angle on I n l e t Reingestion . . . . . . 9-41

. - AIRElPACT CQMFiQNY

x i i i M I ) ( : A4318 LIST OF FIGURES (Continued) T i t l e .

9-34

Effect of Louver Deflection Angle on I n l e t Reingestion . . . . . . 9-42

9- 35 Effect of Louver Deflection Angle on I n l e t Reingestion

. . . . . . 9-43

e 3 6 Effect of N o s e Fan Speed 011 I n l e t Reingestion . . . . . . . . . . .

9-44 9-37 Effect of Nose Fan Speed on I n l e t Reingestion . . . . . . . . . . .

9-45 e 38 Effect of Nose Fan Speed on I n l e t Reingestion . . . . . . . . . . .

9-46 9- 39

Effect of Nozzle Exhaust Splaying on I n l e t Reingestion . . . . . . 9-47

9-40 Effect of Nozzle Exhaust Splaying on I n l e t Reingestion . . . . . .

9-48 9-41 Effect of Nozzle Exhaust Splaying on I n l e t Reingestion

. . . . . . 9-49

9-42 Effect of Nozzle Exhaust Splaying on I n l e t Reingestion . . . . . .

9-50 9-43 Effect of Shielding on I n l e t Reingestion

. . . . . . . . . . . . . 9-52

9-44 Effect of Shielding on Inlet Beingestion . . . . . . . . . . . . .

9-53 9-45 Effect of Shielding on I n l e t Reingestion . . . . . . . . . . . . .

9-54 9-46 Effect of Shielding on I n l e t Reingestion . . . . . . . . . . . . .

9-55 9-47

Effects of Shield Deflection Angle on Ialet Reingestion . . . . . . 9-56

fniet h i n g e s t i o n Sermocoupie fdentifications . . . . . . . . . .

9-48 9-57 Inlet Reingestion Analog Temperature Variations.

9-49 &del Height = 21.0 F t . . . . . . . . . . . . . . . . . . . . . .

9-58 I n l e t Reingestion Analog Temperature.Variation8 . . . . . . . . . .

9-50 9-59 9-51 I n l e t Reingestion Analog Temperature Variations . . . . . . . . . .

9-60 9-52 Left Gas Generator I n l e t Analog Temperatcre Variations.

Model Height = 21 F t . . . . . . . . . . . . . . . . . . . . . . .

9-61 9-53 L e f t Gas Generator Inlet Analog Temperature Variations.

Model Height = 8.3 F t . . . . . . . . . . . . . . . . . . . . . . .

9-62 9-54 Left G a s Generator I n l e t Analog Temperature Variations . . . .

9-63 9-55 Representative Flow Visualization T e s t Results . . . . . . . . . .

9-64 PfDC A4318 Description i n l e t highlight area Noaele exit area Fan inlet capture area Throat area Antisubmarine warfare wing span Butt l i n e Chord Mean aerodynamic chord Drag coef f i c i e n t (D/qS) Thrust calibration coefficient ( F s / ~ A ) / ( F ~ / ~ ~ ) Rolling moment coefficient (!L/qSb) Lateral s t a b i l i t y parameter (aCE/a8) per degree L i f t coefficient (L/qS) Pitching moment coefficient about F/4 (m/qSZ) Yawing moment coefficient (n/qSb)

Directional s t a b i l i t y parameter (ah/ 36) p e r degree

Side force coefficient (Y/qS) Side force parameter (aC,/aS) per degree Drag o r nozzle exit diameter N (lb) o r m ( f t ) Distortion factor ( P T m x - P T a ) /PTAVG Gross thrust a t forward speed Gross thrust of l i f t / c r u i s a unit Gross thrusi; of nose l i f t unit MDC A4318 Description Ideal gross t h r u s t N e t thrust S t a t i c t h r u a t Fuselage Station Acceleration of gravity GE General Electric H We1 height above ground measured from l i f t c r u i s e nozzle & t .

a Rolling montent L Lift L i f t cruise LC LSAE Large Scale Aerodynamics Branch N-m (ft-lb) m Pitching moment McDonnell Alrcraf t Company N-m (ft 5 ) n Yawing moment

Fan speed rn

NF Gas generator fan speed RPM NGG Nozzle t o t a l pressure r a t i o NPR S t a t i c Pressure P Ambient static pressure P A Total pressure PT Freestream dynamic iJzegF'?,?= !1./2~0V0 R Radius Reference Center RC Corrected fan r o t a t i o n a l speed Wing area O K ( O R ) Total temperature

---

Thermocouple A-RAFT COMPLINV mi

mC A4318

Descriptim

S e S L

TEL/TEB Trailing edge l e f t / r i g h t

-

TBu/TED Traillng edge up/down :et t o t a l temperature O C (OR) *J ;et Valocity dsec (f t/sec) 0, Velocity at the throat d m c (ft/sec) VTH F Zeestream velocity d a e c (ft/sec) VO

-

cpo/V,> st velocity ratio

m Vertical takeoff

V/STOL Vertical/short takeoff and landing W Airflow rate W.L. Waterline Windmilling W M Y Side force GREEK SYMBOLS Angle of attack AngZe of s i d e s l i p Incremental

Inlet temperatne rise ( T T ~ - T T ~ )

Jet tempe-sture rise (TJ - TTO)

Aileron deflection (positive I s TED) ',eft aileron def l e c t l o a Right aileron deflection Relative static pressure (P/14.696) Pls? keflection (positive I s TED) Horizontal t a l l deflection (positive is TED) L i f t cruise unit geometric deflection Description Nose l i f t unit geometric deflection Rudder deflection (positive is TEL) Inlet shield deflection angle Relative t o t a l pressure (P~/14.696) Nozzle yaw vane deflection (positive TEL) Resultant thrust vector angle Lift cruise unit thrust deflection angle Nose l i f t unit thrust deflection angle Relative t o t a l temperature (-/518.7) Sweep angle Freestream density Freestream Inlet throat Fan entrance 3 Fan exit MDC A4318 1. INTRODUCTION 1.1 BAamuIulj McDoanell Aircraft Company (MCAIR), a division of McDonnell Douglas Corpora- tion, over the past several years has conducted a comprehensive analysis of V/STOL a i r c r a f t designs.

O n e of the configurations analyzed, a l i f t / c r u i s e fan multi- mission a l r c r a f t with interconnected propulsion units, provides s i p i f i c a n t advancemeute in V/STOL aircraft performance and operational capabilities. Large improvemants are provided in payload/range, speed, a l t i t u d e , safety , r e l i a b i l i t y and maintainabflity which are not presently achievable i n e i t h e r rotary wing o r other proposed V/STOL vehicles. This iatetconnected l f f t / c r u l s e fan concept can s a t i s f y the military needs f o r V/STOL multimission a i r c r a f t as w e l l as many civil needs f o r u t i l i t y a i r c r a f t f o r support of construction, lumberiag, o i l exploration o r developmeat sites located in areas d i f f i c u l t to reach rapidly by other modes of transportation.

The MCAIR l i f t / c r u i s e fan a i r c r a f t is a fimted wing vehicle powered by three i d e n t i c a l turbotip driven fans t h a t are pneumatically interconnected t o each other aud t o two (or three) gas generators (depending on the s p e c i f i c a i r c r a f t design) during the powered l i f t f l i g h t mode. During cruise mode operation the gas sen- erators power only the two over-the-whg l i f t / c r u i s e fans. This unique applica- tion of f l i g h t proven l i f t fans provides a variable bypass r a t i o propulsion system which maximizes l i f t and control during VTO, provides excellent short takeoff characteristics, allows e f f i c i e n t engine matching for cruise and l o i t e r , and provides f o r retention of symmetrical thrust, good control margins and v e r t i c a l landing capability during one-engine-out operation.

Recognizing the need for development of technology f o r such V/STOL m u l t i - mission a i r c r a f t f o r Naval purposes that also may serve many future c i v l l - u t i l i t y a i r c r a f t requirements, the N A S A Ames Research Center, with Navy support, contracted with MCAIR t o design a large scale powered model f o r test i n t h e NASA A m e s 40 foot by 80 foot wind tunnel. The i n i t i a l design work w a s performed under Amendment No. 19 t o Contract NAS2-5499. The remainder of the e f f o r t was done under Contract NAS2-8655.

The model was t o be approximately 0 . 7 scale of a potential Navy multimission a i r c r a f t and was t o u t i l i z e fans and gas generators already existing in the N A S A test hardware inventory. In addition, the contract covered engineering l i a i s o n support during fabrication of the model in the NASA A m e s shops, engineering and technical support during the active test phases, data reduction MDC A4318 and a n a l y s i s , and r e p o r t i n g of t h e results. This document is t h e f i n a l r e p o r t of t h e s e a c t i v i t i e s . I n a d d i t i o n , a NASA d a t a r e p o r t (Reference 1) w a s p u b l i s h e d s h o r t l y a f t e r completion of t h e tests., 1.2 4 0 ' X 80' WIND TUXNEL TESTS The main o b j e c t i v e s of t h e tests i n t h e 4 0 ' x 80' wind t u n n e l included assessment of: o Powered f l i g h t c h a r a c t e r i s t i c s , o Aerodynamic f l i g h t c h a r a c t e r i s t i c s , power on and o f f , t a i l on and o f f , and Propulsion a i r induction system c h a r a c t e r i s t i c s . o The powered l i f t test d a t a , from z e r o speed through t r a n s i t i o n , were used t o e v a l u a t e f l i g h t procedures and v e c t o r i n g schedules d u r i n g t r a n s i t i o n and t o estab- l i s h aerodynamic-propulsion induced e f f e c t s and flow f i e l d s w i t h v a r y i n g f a n t h r u s t , t h r u s t v e c t o r angles, jet v e l o c i t y ratios, and combined powered and aerodynamic c o n t r o l inputs. Longitudinal and l a t e r a l - d i r e c t i o n a l C h a r a c t e r i s t i c s i n t h e l o i t e r and c r u i s e f l i g h t modes were e s t a b l i s h e d from t h e aerodynamic test data.

test d a t a were used t o e s t a b l i s h d i s t o r t i o n p r o f i l e s and The p r o p u l s i o n system i n l e t performance throughout t h e test o p e r a t i n g range. T e s t v a r i a b l e s included angle of a t t a c k , s i d e s l i p angle, fan t h r u s t , t h r u s t v e c t o r a n g l e s , tunnel a i r f l o w v e l o c i t y , aerodynamic c o n t r o l d e f l e c t i o n s , and simulated powered l i f t c o n t r o l inputs.

1 . 3 OUTSIDE STATIC TESTS The o u t s i d e s t a t i c test program included ground e f f e c t s t e s t i n g and p r o p u l s i o n system c a l i b r a t i o n tests. The ground e f f e c t s test program included tests a t 21.0, 8.3, and 3.3 f o o t h e i g h t s and measured t h e t o t a l i n s t a l l e d l i f t loss and i n l e t re!L.6estion c h a r a c t e r i s t i c s of t h e model. The c a l i b r a t i o n tests included t h e :..ner- :on of a f a n performance map and tests of each l i f t u n i t f o r comparison w i t h t h c s t a t i c c a l i b r a t i o n s done i n t h e 40' x 80' tunnel. Flow f i e l d v i s u a l i z a - t i o n tests were a l s o conducted u t i l i z i n g smoke. Force and moment d a t a , f a n and gas generator performance d a t a , and i n l e t performance d a t a were recorded d u r i n g the te..t program.

MCDOUURLL AIRCRAFT COMPANV 1-2 MDC A4318 2. MODEL DESCRIPTION The aircraft model t e s t e d i n t h i a program w a s a l a r g e scale (approximately 70%) powered model of a subsonic fixed wing lift/cruise f a n V/STOL a i r c r a f t concept configured o r i g i n a l l y f o r t h e Navy ASW missioa. This configuration w a s the result of several design compromises required to provide multimission a d a p t a b i l i t y f o r operational usage aboard the Navp's VSS, LBP, LEA, and DD 963. Aerodynamic refine- ments of this design were based on s m a l l scale model low speed and high speed wind tunnel test data. These data were incorporated i n t o a mathematically accurate d e f i n i t i o n of the complete aircraft mold line surface through use of the MCBIB Computer Aided Design Drafting ( O D ) i n t e r a c t i v e graphics system. While the standard construction functions were used t o initiate this e f f o r t , it was refined through extensive use of t h e Parametric Cubic (PC) curve and PC patch routines t o a r r l v e at P C surface d e f i n i t i o n .

For the model design the f u l l scale mold line d e f i n i t i o n was reduced t o 70X of f u l l scale. Model construction of a tubular truss type s t r u c t . -e with l o c a l rings supporting t h e combination metal and f i b e r g l a s s mold line skins w a s s e l e c t e d i n order to reduce c o s t s by allowing r e l a t i v e l y loose tolerances of the s t r u c t u r e with reasonably accurate control of the external mold l i n e shapes. Cuts were made a t appropriate s t a t i o n s as determined by the model layout drawings f o r headers, r i b s , longerons, etc.; d e t a i l s and reference line information were added, and drawings were hard-copied on s t a b l e Mylar. Part number c a l l o u t s , hole sizes, and p a r t s l i s t were added manually along with small d e t a i l c u t s t o obtain the l e v e l of informatioo required f o r hardware manufacture and model assembly a t NASA Ames.

The main f e a t u r e s of the model included three gas generator driven turbotip fans, and variable geometry f o r all c o n t r o l surfaces and vectoring system components.

A photo of the model is shown i n Figure 2-1. The s i z e and d e t a i l design of the model were based on u t i l i z i n g e x i s t i n g propulsion system components, including the

gas generators , turbotip fans, and vectoring sys tem components supplied by

NASA h a . The physical s i z e and performance c h a r a c t e r i s t i c s of the TS8-GE-8B $as generator and low pressure r a t i o (1.08) GE-X376B turbotip f a n were the predominant considerations i n s i z i n g t h e model. Based on performance estimates of t h i s system, a model scale was selected t o provide i n l e t velocity r a t i o s (VO/VTH) t y p i c a l of f u l l scale higher prssaure r a t i o fan systems during jet velocity r a t i o (V0/Vj) excursions. A schematic of the model i l l u s t r a t i n g the major propulsion syetem components is shown in Figure 2-2. The model had an o v e r a l l length of 10.26 meters (33.7 f t ) , a span of 8.68 meters (28.5 f t ) and a height of 2.76 meters MOCA4318 FIGURE 2-2 LARGE SCALE POWERED MODEL PROPULSION SYSTEM LWCruise Fan Inlet X3768 Forward Fan T6888 Forward Engine Inlet DumingPlenum Lift/Cruise Engine Inlet LiWCruise Nozzles 738-88 LiWCruiseEngine X376B Lift/Cruise Fans M e R L L A-RAPT OOMPANV 2-3 MDC A4318 The b k f c geometry and overall dimensions of the model configuration (11.9 f t ) .

Descriptions of the airframe and propulsion system com- are shown in Fi'gure 2-3.

ponents are presented below.

-- - -

2.1 FUSELAGE The basic fuselage shape of the a i r c r a f t accommodates side-by-side seating i n the forward fuselage, and provides the necessary volume in the center fuselage f o r satisfying the needs of the multimission role. This wide bodied design allows f o r the i n s t a l l a t i o n of the l i f t fan unit in the forward fuselage section of the a i r c r a f t .

The fuselage section of the test model contained the main support truss, which The three gas in turn supported the wings, tdl section, and turbotip fan units.

generators, interconnect ducting, f u e l distribution l i n e s , lubrication system, instrumantation, f i r e extinguisher system and forced air cooling l i n e s were a l l housed within the fuselage. Vent louvers were i n s t a l l e d on the lower and upper surfaces of the fuselage t o provide i n t e r n a l cooling air f o r the test model.

2.2 WING The design concept incorporated a low wing with the lower surface flush with the bottom of the center fuselage. The basic wing had an aspect r a t i o of 4.5, a taper r a t i o of 0.30, and a quarter chord l i n e sweep of 25'.

Total wing planform area was 16.75 m2 (180.3 f t 2 ) .

Further d e t a i l s of the wing geometry are given i n __ --_I- Figure 9 - 3 .

The basic wing had different a i r f o i l sections inboard and outboard of the lift/crulse fan nacelle/wing intersection. The Inboard wing panel had an NACA 4416 a i r f o i l section a t B . L . 37.8, and the outboard wing panel used modified supercritical a i r f o i l sections a t root and t i p . The transitions between these three specified s t a t i o n s were s t r a i g h t l l n e elements. Coordinates of the wing a i r f o i l sections are presented i n Figure 2-4.

The coni& surfaces consisted of p l a i n f l a p s and ailerons both hinged a t The flaps had s p e c i f i c deflection angles of O', 15', 30' the 75% chord line.

The ailerons had a deflection range of and 45" and were manually positioned.

- + 25' and were remotely operated.

_- 2.3 EMPENNAGE The a i r c r a f t empennage consisted of a "T" t a i l configuration with a movable Iiorizontal sttrbilatot and fixed v e r t i c a l s t a b i l i z e r . The geometry of the v e r t i c a l and horizontal tail is presented i n Figure 2-3.

Both the v e r t i c a l and horizostal t a i l components were detachable from the model for tall-off testing.

MDC A431 8 I 4 MDC A431 8 FIGURE 2 4 LARGE SCALE LIFT/CRUISE FAN AIRCRAFT MODEL Wing Airfoil Ordinates _ _ _ _ _ - - - - __ . . . . .

- - _ _ a 5 b - b

Wing station (0.4422) Theoretical ip (0.941 5)

Station Exposed R o o t (0.221 5)

Modified, Mod J, NACA 8416 Supercritical, t * 14% Supercriti l,t=8%C

-

YL, %c Yu, %c YL. %c vu, 9Bc vL. %c

- -

~ ~- -

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

-1.435

1.26 2.471 - ?.467 3.275 -1 .w 1.422

-1.840 2.6 3.233 -3.21 8 4.448 -2.645 1.836 -2.334 5.0 4.126 4.069 6.123 -3.486 2.307 4.653 7.371 -3.957 2.641 -2.678 7.6 4.729 10.0 5.208 4.099 8.363 -4.245 2.902 -2.944 -3.335 16.0 5.945 -5.756 9.888 -4.459 3.293 -6.1 77 10.933 -4.427 -3.598 20.0 6.481 3.569 25.0 6.878 -6.433 11.648 -4.245 3.?68 -3.766 -3.682 30.n 7.1 65 -6.560 12.000 -4.000 3.904 -3.837 40.0 7.478 -6.. 5 12.000 -3.467 4.029 1 1.232 -3.531 50.0 7.484 -5.753 -2.901 3.988 4.503 9.920 -2 283 3.785 -2.766 80.0 7.229 70.0 -2.786 8.139 -1.653 3.420 -1.465 6.662 -1.091 5.920 -1.og9 2.829 4.084 80.0 5.685 90.0 -0.194 3.286 -0.608 1.843 9.496 3.980 96.0 2.61 6 -0.210 1.781 -0.384 1 .ogo j J.342 100.0 -0 447 0.1 71 -0.171 0.108 0.214 0.496 2.822 1.154

LE Radius, 96c 3.041

0.891 Chord Length, m 2.409 2.049 ft (2.922 1 (8.231 1 (6.7231 -2.77 Incidence, deg 3.23 2.59 Mn: A4318 The vertical rail utilized a b a s i c NACA 658010 a i r f o i l section and w a s of oo, The rudder had mmlnal deflection equipped w i t h a moweable d e r .

+loo, too, and +so, and waa luulually positioned.

The horizontal tail utilized o syumetrical W C A 64AOXX a i r f o i l section w i t h The horizontal tail a thickness ratlo of 0 . 1 0 a t the root and 0.08 at the tip.

was a -piece unit that was r-tely operated, and had a d e f l e c t i o n range of - +20° .

2.4 A I R IHDucTIolo spsran The model air ind *ion systeas ccasisted of the l i f t / c r u i s e fan inlets, the nose fan inlet, and gas generator fnlets for each type fan installation as s h m in Figure 2-3. A description of each inlet i s presented below.

LiftICruise Fan Inlets The l i f t / c r d s e fan inlets were located over the wing and adjacent t o the fuse- lap in a fully integrated design coacept. They w e r e fixed geometry inlets with an

internal contraction ratio (b/b) of 1.25. They had a 2:l e l l i p t i c a l l i p pro--

file and cubic duct contours internally. A low drag modified e l l i p t i c a l cowl con- tour w a s used externally. The d e t a i l e d geometry of t h e liftlcruise inlets is presented in Figure 2-5.

L i f t / C r u i s e Generator I n l e t The l i f t / c r u i s e engine inlets w e r e s i d e mounted with fixed geometry.

They also had an inlet contraction r a t i o of 1.25, a 2:l e l l i p t i c a l l i p shape i n t e r n a l l y , For ease in fabrica- and l o w drag, modified e l l i p t i c a l cowl contours externally.

Parabolic tion, a straight-line-eleamnt conical duct was u t i l i z e d internally.

boundary layer d i v e r t e r s were incorporated between the inlets and the fuselage.

Details Details of the gas generator inlet geometry are presented i n Figure 2-6.

of the shields added as a means of reducing hot gas reingestion during operation a t l o w ground heights are presented in Figure 2-7.

. - Nose L i f t ~ a n Inlet The nose fan inlet was located i n the nose of the aircraft j u s t downstream of the radome and forward of the cancpy. It was a flush mounted i n l e t with an o v e r a l l resultant contraction r a t i o ( A ~ / A F ~ ) of 2.09. The forward s e c t i o n of the i n l e t ' had a l i p thickness r a t i o ( Y / k ) of 0.30. s decreased t o a minimum thickness ratio of 0.20 a t the s i d e s , and remained a constant over %fZ .ection of the i n l e t . I n l e t turning vanes wsre not incotpori\ted f r @ 1 . . i design. It had ?astad t o a 2 : l e l l i p - a 1.4:l elliptical l i p p r o f i l e a t the leadin8 '.:y

tical prof€le lit the side. The 2:l p r o f i l e w& . ' I . : ever the a f t section of

the i n l e t . I n l e t closure door6 were not included on t h i s test model. F o r cruise - A B A W - A I v 2- 7 MDCA4318 MDCA4318

1 %

-

rua T : t MOCA4318 A B A F T 2-10 w)C A4318 mode t e s t i n g (nose fan o f f ) , an Inlet closure panel w a s I n s t a l l e d on t h e Inlet.

Ltafls of the.nose l i f t fan Inlet geometry are presented In Flgure 2-8.

b e Pan Gas Generator Inlet The nose faa gas generator inlet design consisted of two flush mounted inlets, each ducted i n t o a conmon plenum located upstream of t h e a f t fa- gas generator.

The inlets were located on the upper surface of t h e lift/cruise gas generator nacelles at the approximate wing leading edge s t a t i o n . Each inlet had a cmtrac- tlon ratlo (&/Am) of 4.0 and a d i f f u s i o n r a t i o (!.:n/b) i n t o t h e plenum of 1.5. A l l Inlet U p 8haping consisted of 2:l e l l i p t i c a l p r o f i l e s wtth vcrryln8 l i p thickness around t h e periphery. For cruise mode testing, Inlet cloctrre panela were Installed on each inlet. Details of t h e nose fan gas generator I n l e t geometry are pe-ted In Figure 2-9.

Three i d e n t i c a l 36" diameter GE-X376B turbotip fans w e r e i n s t a l l e d In t h e model, each one driven by a modified T58-GE-8B gas generator. The engine and fan f o r each l i f t u n i t were interconnected with steel ducts and bellows arranged as shown in Plgure 2-2. T5e gas generators, turbotip fans, fntercoanect ducting and bellows were exlssting hardware itenis supplied by t h e Large Scale Aerodynamlcs Branch (LSAB) of W A / A m e s . The design characteristics of the gas generator and fan are presented In Figure 2-10. The performance values given are GE engine s p e c i f i c a t i o n values, and are presented f o r reference only.

2.6 TIIRaST VECTORING SYSTEM The exhaust nozzle t h r u s t vectoring systems u t i l i z e d on t h e nose l i f t unit and the two l i f t / c n t i s e units were eltisting hardware components previously used and supplied by t h e LSAE of NASA/kmes. Descriptions of the l i f t / c r u i s e vectored nozzles and the nose vectoring louvers are given in Figures 2-11 and 2-12, respectively, and are discussed below.

Lift/Cruise Vectored Nozzle The l i f t / c r u i s e vectored nozzle consisted of a fan exit d i f f u s e r , t h r u s t vec- t c r i n g hood segments, and a nozzle eKit cone, as shown in Figure 2-ll. The fan exit d i f f u s e r had a diffusion r a t i o (AEOO~/AFAN EXIT) of 1.50 and was used f o r both (cruise mode) test arrangements. Thrust vectoring w a s the vectored and noavectored achieved with t h e fixed diameter detachable angular hood segments, so arranged as t o provlde geometric deflection angles ( 6 ~ c ) of 2 3 O , 38", 5 6 O , 71' and 90'. The I n i t i a l 23" hood segment w a s equipped w i t h detachable cover panels on the upper surface, providing exhaust ports f o r simulating t h r u s t reduction modulation. The AIRCCILIC'I ColwplrrNy 2-11 MDCA4318

E

e Q a U

a

b

MOCA4318

F W R E z*

FORWARD ENGINE INLET DESIGN GEOMETRY

Contraction Cone 1

Inlet Amas Highlight Area (AHL) = 2 0 0 In?

Throet Area (Am) 40.6

Duct Area (AD) = 74.5 i

Plenum Area CAP) = 176.7 Ellipitical Top vim Profiles Side View A W A C T COMCANV 2-13 FIGURE 2-10 GAS GENERATOR AND NRBOTIP FAN DESIGN CHARACTERISTICS ___ Paint lbfomme ~IntermedW Power)

Air Flow ...................................... 12.4 lblsec

Compressor plassure Ratio ........................ &0:1

Turbine fnfet Temparature ........................ 171@F

Exhaust Gas Temperatuta. ........................ 12s0°F

&ginaspeed ............................ . W O O rpm

42 in. Exit Die - 1

f----------

Design Foim petfonnance (100% Speed)

Air Flow ...................... .. ............... 153 I b h c

Specific Flow .................................. 27 Ib/mc.ft2

Fan Pressure Ratio .............................. 1.08

Admission Arc.. ................................ 180'

Fan Speed ( 100%) ............................... 4074 rpm

A-RAFT CoMp19NV 2-14 MDC A431 8 FIGURE 2-11 LIFWCRUISE UNI~V%TORING SYSTEM GEOMETRY Thrust Modulation Ports d - - - - - - *

--------

Two 10% Thick, 12 in. Chord Articulated Yaw Vanes I I I ~ c

r Diff user

. .

--------

-LL A W A C T COMCLINV 2-15 MOCA4318 FIGURE 2-12

NOSE LIFT UNIT VECTORING systm GEOMETRY

Louver Actuator

Vectoring Louvers -

b . 4 \\ Fixed beding Edge

Hinge tine _ _

( 2 6 m h o d TWO 10% midk, 12 in. mod Articulated Yaw Vanes L I B A C T eOMCYINY 2-16 MDC A4318 thrust vectoringhood segments had constant area turning with a turning radius of 0'.54 R/D. The:nozele exit cone w a s a l s o detachable, and w a s equipped with two 10Z thick, manually positioned, a r t i c u l a t e d yaw vanes. These vanes provided lateral

vectoring of O o , 5 ' and 2 1 3 . O t o produce yawing moments. The nozzle cone had a

fixed nozzle exit area of 0.7677 m2 (1190 in.2) with an exit contraction r a t i o (AHOOD/ANOZ) of 1.16.

inch long constant area nozzle duct ( 6 ~ c = 0'1, a 17 For cruise mode t a s t i n g extension w a s i n s t a l l e d downstream of the diffuser section. The c d s e nozzle etdt cone, which w a s attached t o t h i s extension, did not include exit yaw vanes. The cruise nozzle exit area w a s .6935 m2 (1075 i n . 2 ) . An a l t e r n a t e cruise nozzle exit area of 0.6 n ? (930 inD2) was provided by a nozzle ring which attached t o the exit cone. The cruise and vectored U f t / c r u i s e nozzles are shown in Figure 2-11.

Nose L i f t Unit Vectored Nozzle The nose l i f t u n i t t h r u s t vectoring nozzle system u t i l i z e d an existing, Fourteen l o w camber louvers, each with remotely activated louver and drive system.

a thickness r a t i o of lox, provided thrust vectoring over a range from 105' t o 3 0 ' .

T w o 1 O X thick, manually positioned, articulated yaw vanes located beneath the lou-

vers provided yaw vectoring of O', 5' and - +12". The yaw vanes were of the same

design as those on the l i f t / c r u i s e units, and were detachable from the model. For cruise mode testing, the complete louver and yaw vane vectoring system w a s remwed and a lower surface nozzle u n i t cover panel was i n s t a l l e d on the model. Details of the nose l i f t unit vectoring system are presented in Figure 2-12.

MDC A4318 3. MODEL INSTRUMENTATION The l a r g e ' s c a l e powered model was f u l l y instrumented with pressure and tem- perature pickups located at various positions on the airframe and propulsion system compments. The quantity, type, and location of the instrumentation f o r each model component are described in the sections that follow.

3.1 AIRFRAME INSTRUMENTATION A t o t a l of 105 static pressuze ports were installed on the external surfaces The locations of the pressure.

of the wing and fuselage sections of the airframe.

ports on each component are described below. N o data from t h i s instrumsntation are included i n t h i s report due t o problems encountered i n the reduction process.

wfng A t o t a l of 78 static pressure ports were installed on the l e f t wing, aileron, and flap surfaces, distributed along b u t t l i n e s located a t four spanwise positions.

lower l e f t wing surfaces were instrumented. The locations of Both the upper and a l l Wing pressure ports are presented in Figure 3-1.

Forward Fuselage A t o t a l of 27 static pressure ports were installed on the forward fuselage of the model. This t o t a l included 16 ports 0- the l e f t side of the nose section and 1 1 on the lower fuselage surface j u s t downstream of the forward fan exhaust.

The detailed locations are shown i n Figure 3-2.

- i peanage Flow Survey Rake

Dynamic pressure and flow direction were measured at the pivot axis of the horizontal tail plane by directional p i t o t static probes a t the following six butt lines: 29.0, 42.1, 55.3, 68.4, 81.5, and 94.6.

3.2 PROPULSION SYSTEM INSTRUbENTATION The propulsion system instrumentatiob included static pressure, t o t a l pressure, and t o t a l temperature measurements a t various locations on the components of the three l i f t units. The components of the propulsion system that were Instrumented included the following: o Left l i f t / c r u i s e fan and gas generator i n l e t s o Nose fan and gas generator i n l e t s o Left and nose gas generator exits o Left, right, and nose fan and t i p turbine e x i t s o Left and nose fan nozzle exits.

MDC A431 8 FIGURE 3-1 LOCATION OF WING SURFACE STATIC PRESSURE PORTS (Left Wing Only)

._ - - -. -- - _ _ . - __ - - ._ - __ - ---

-

- BL-81-20' BL 161.0, BL 106.0' % Chord1 UPPU Surface

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- -

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40.13 90.0 70.20 59.52 74.10 74.10 42.36 42.36 96.0 62.82 62.82 44.14 99.0 77.22 77.22 44.14 66.47 66.47

-

100.0 78.00 44.59 66.13

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QP76-0022-204 k1obO1vllRILI. A B R A C T COlUCMNV 3- 2 MDC A431 8 FlriWRE 3-2 LOCATION OF F-Ol3WAl?D-FU’3EJLA(3€ SfA’iIC PRESSURE -_ PORTS.

Lw. Hand @de MK: A4318 A t o t a l af 108 static pressure ports, 247 total pressure probes, and 79 t o t a l

__

remperature probes were Installed In the, propulsion systea of tte model.

A tabla sumarigiag the propulsion system instrumsatation and the specific purpose f o r each inst-tion is presentad in F l y r a 3-3.

A description of the quantity, type, and location of the instr\wertatioa for each propulsion system component is picsented below.

Fan Inlets The left lfft/cruise aud nose l i f t fan Inlets w e r e each iastrursented with inlet surface static pressure ports, located at warlous positions on the internal and external surfaces. The l e f t lift!cntise nacelle included a t o t a l of 22 static pressure ports located on the upper and side i n l e t lip. the lower Internal duct surface, and on the external upper uacelle surface. In addition, six static pressure ports were also installed on me upper wing surface ahead of the inlet.

The locations of these and the inlet ports are shown I n Figure 3-4. The locations of the nose let fan i n l e t static pressure ports are shown i n Figure 3-5. A total of 10 ststjc pressure ports were installed on the nose and side inlet l i p s .

Engine Inlets The l e f t l i f t l c r u i s e and nose fan engiie inlets were also hstrumented with static pressure poics on t h e surfaces of the inlets. A t o t a l of 10 Inlet l i p static pressure ports were installed on the internal surface of the l e f t l i f t / c r u i s e The locations are shown i n Figure 3 - 6 . Five engine inlet side and lower lips.

inboard l i p static pressure ports were installed 021 the l e f t nose fan engine i n l e t , This l e f t iFht also included an eight-probe t o t a l tem- as shown in Figure 3-7.

perature rake t o measure raingestion temperature levels a t the inlet throat station.

locations of the thermocouples are also presented in Figure 3-7.

The Fan Face The l e f t l i f t / c r u i s e , ' a n face and nose l i f t fan face w e r e both Instrumented with Identical inlet performance rakes. An 8 leg, 48 t o t a l pressure probe rake t o measure i n l e t performance, along with a 4 leg, 8 t o t a l temperature probe rake t o measure reingestioq temperatures, was installed a t the fan face, j u s t upstream of the X376B forward support frame on each l i f t unit. Both the pressure and tempera- w a l l s t a t i c pressure ports were ture rakes were equal-area-weighted rakes. Four also located at the rake s t a t i o n of each inlet. The Instrunentation locations for the fan face rakes a r e presented in Figure 3-8.

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Front view Left s i 5 Lower Lip Statics at BL 33.5 Highlight Static Pressure Locations Throat 0.35 / - 2.05 HighlightJ -LL A-RAPT COWMNV 3-8 MDCA4318 FIGURE 3-7 FORWARD ENGINE INLET INSTRUMENTATION Left Side Only 5 Inboard Lip Statics 8 Total Temperaarm Robes at Inlet Throat FS 240.6 4 . 2 5 i n . -

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6P70-0622.7 Cut at FS 240.6 MDC A431 8 FIGURE 3-8 FAN FACE INSTRUMENTATION 8 Leg/48Probe Total Pressure Rdce R - = 8.1 in.

Rtm - 18Oin.

14.72 13.17 11.43 6 9 . 3 6 Inlet Rake Locations A-ACT -NV 3- 19 MDC A4318 A l l t o t a l p r e s s u r e and t o t a l temperature rake probes i n s t a l l e d i n t h e model bere made of lt8" O . D . x .032" w a l l t h i c k n e s s stainless tee1 tubing. The p r e s s u r e probe i n l e t s were chamfered a t 15'.

A l l thermocouples i n s t a l l e d in t h e model were made of 30 gauge i r o n / c o n s t a n t a n wire.

A t e f l o n l i n e r w a s u t i l i z e d w i t h i n t h e steel t u b i n g f o r i n s u l a t i o n .

Engine Face l e f t l i f t / c r u l s e engine face and t h e nose f a n engine f a c e w e r e both instru- The mented w i t h inlet performance rakes. A 4 l e g , 16 t o t a l p r e s s u r e probe r a k e t o measure Inlet performance, along w i t h a 4 l e g , 8 t o t a l temperature probe rake t o measure r e i n g e s t i o n temperature l e v e l s , w a s i n s t a l l e d j u s t upstream of t h e engine f a c e on each unit. Two w a l l static p o r t s were a l s o l o c a t e d a t t h e r a k e s t a t i o n of each engine. The i n s t r u m e n t a t i o n l o c a t i o n s f o r t h e s e engine f a c e rakes are presented in F i g u r e 3-9.

Fan and T i p Turbine Exit A l l t h r e e t u r b o t i p f a n units on t h e model were instrumented i n like manner a t t h e f a n and t i p t u r b i n e s t a t o r exits. The fan exit i n s t r u m e n t a t i o n c o n s i s t e d of a 6 l e g 30 probe t o t a l p r e s s u r e rake, a 3 l e g 9 probe t o t a l temperature rake, and 12 exit static p r e s s u r e p o r t s , 6 each on t h e hub and o u t e r w a l l . The t i p t u r b i n e exit i n s t r u m e n t a t i o n c o n s i s t e d of 4 equal-area-weighted t o t a l p r e s s u r e probes, 4 t o t a l temperature probes, and 5 o u t e r w a l l static p r e s s u r e ports. The purpose of t h e f a n and t i p t u r b i n e exit i n s t r u m e n t a t i o n w a s t o measure t h e b a s i c performance of t h e t u r b o t i p f a n s i n c l u d i n g t h e a i r f l o w , t h r u s t s , and jet v e l o c i t i e s . The d e t a i l e d l o c a t i o n of t h e f a n and t i p t u r b i n e exlt i n s t r u m e n t a t i o n is p r e s e n t e d i n F i g u r e 3-10. A c r o s s - s e c t i o n a l drawing of t h e t u r b o t i p fan i l l u s t r a t i n g t h e p o s i t i o n i n g of both t h e inlet and f a n exit r a k e s is p r e s e n t e d i n F i g u r e 3-11. Details of t h e t o t a l p r e s s u r e and temperature probes are a l s o shown.

Ennine Exit The l e f t l i f t / c r u i s e and nose f a n engine exhaust d u c t s were each instrumented w i t h t h r e e w a l l static p r e s s u r e p o r t s , l o c a t e d approximately one duct diameter downstream of t h e engine exit. The p r e s s u r e p o r t l o c a t i o n s are shown i n Figure 3-12. In a d d i t i o n t o t h e s e static p r e s s u r e measurements, t h e eight-probe engine EGT harness was u t i l i z e d on a l l t h r e e engines t o measure t h e engine exhaust gas t o t a l temperatures.

MDC A431 8 FIGURE 3-9 ENGINE FACE INSTRUMENTATION 2 Wall Statics at 1-80' Spacing

- LIH Inlet (Top/BottomJ

- Forward I,-!at (Sides)

Total Pressure Rake 8 Total Temperature Probes a t 90° Spacmg Thermocouple Locations Pressure Tube Locations Radius (in.)

Pro; No. - 1

TIC No.

5.20 4 9 9 4.40 4.32 3.69 3.62 4 3.14 3.04

Inlet Rake LocationsJ \ ' -

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~ ~ 7 0 oeaz i o MCDONNELL A I ~ C ~ A FT COMPANY 2-iL MDC A431 8 FIGURE 3-10 FAN AND TIP TURBlNE EXIT INSTRUMENTATION Rear View Left Side Pressure Tube Locations Thermocouple Locations Radius (in.)

14.78 Exit Rake Locations GP76 0622 2 MCQONNELL AIRCRAFT COMPE.NY 3-13 MDC A431 8 FIGURE 3-7 1 FAN INLET AND EXIT INSTRUMENTATION INSTALLATION Left LiftKruise and Nose Fan Units Turbine Exit Statics 7 Rake Station Rake Station Turbine Exit Probes Nose Fan Inlet Lip ‘ , I --- Outer Wall Statics

- -

LiftKruise Inactive Arc Fan Duct Rotor I ‘ Fan Face Rake

~ 1 +-- Fan Exit Rake

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Exit Stators LiftKruise Fan Hub Exit Hub Hub Statics , Nose Fan Hub . .

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4- -i MCOONNELL AIRCRAFT COMPd NY 3-1G MDC A431 8 FIGURE 3-12 ENGINE EXIT INSTRUMENTATION Bottom Exit Locations GP76 5622 9 MCDONNELL AIRCRAFT COMPANY 3-13 MDC A4318 Nozzle Exits The left liftlcruise and nose fan lift units were each instrumented at the nozzle exit for the purpose of assessing the nozzle exit flow profiles. The left liftlcruise nozzle exit included 10 rota1 pressure probes attached to the leading edge of the two fixed yaw vane struts, and 4 external nozzle exit base pressure static ports. The nose fan nozzle exit included 7 total pressure probes attached to the leading edge of the louver centerline drive strut, and canted 1 5 ' to align with the fan exit. The details of the nozzle exit instrumentation are presented for both units in Figure 3-13.

MCDONNELL AIRCRAFT COMPANY 3- 16 MDCA4318 FIGURE 3-13 NOZZLE EXIT INSTRUMENTATION Nose Lift Unit Pressure Probes \ / I+.- I - Louvers-

--

Yaw Vane 4 ‘7 5.0 in. (Typ)

Left LiftKruise Unit MCWONNELL AIRCRAF T COMPANY 3- 17 MDC A4318 4. W I K D TUNSEL-TEST FACILITY T h e model was t e s t e d i n t h e NASA/Ames Research Center 4 0 ' x 80' wind tunnel.

Tho tunnel i s a low SFecd closed c i r c u i t continuous flow t u n n e l w i t h a speed range from 0 to 180 knots which o p e r a t e s a t a n e a r c o n s t a n t t u n n e l t o t a l p r e s s u r e o f one

atmosphere and provides Reynolds numbers up t o 2 x lo6 p e r f o o t . The test section

is 40 f e e t high, 80 f e e t wide a t t n e t u n n e l c e n t e r l i n e , and 80 f e e t long. The c e i l i n g cf t h e tunnel t e s t s e c t i o n opens t o provide f o r model i n s t a l l a t i o n . The wind t u n n e l model bslar.c -s a six-component scale type balance, l o c a t e d d i i e c t l y below t h e test s e c t i o n . The uind t u c n e l c o n t r o l room is l o c a t e d a d j a c e n t t o t h e test s e c t i o n . A c r o s s s e c t i o n drawing of t h e t u n n e l test s e c t i o n and a d j a c e n t F u r t h e r d e : a i l s of t h e 40' x 80' wind t u n n e l are areas is shown i n Figure 4-1.

contained i n t h e u s e r ' s guide of Reference (2).

4.1 TEST ARRATACENEXT The model w a s supported on t h e t u n n e l balance w i t h t h r e e support s t r u t s , two a t t a c h e d t o t h e main wing s p a r and one a t t a c h e d t o t h e a f t fuselage. A l l model l e a d s xere a t t a c h e d t o t h e t v o main support struts which were f i x e d i n height.

The t a i l s t r u t w a s remotely v a r i e d i n h e i g h t t o change a n g l e of a t t a c k . The model w a s p o s i t i o n e d a t t h e c e n t e r l i n e of t h e t u n n e l , 20 f e e t above t h e t u n n e l f l o o r .

Photos of t h e model i n s t a l l e d i n t h e test s e c t i o n are presented i n F i g u r e s 4-2 through 4-6.

The w d e l was equipped with s e v e r a l monitoring and c o n t r o l systems. Leads from t h e s e s y s t c ~5 were a t t a c h e d t o t h e main support s t r u t s and routed down through t h e balance room and up t o t h e main tunnel c o n t r o l room. The model sstbsystems wich external leads included t n e following: o Fuel system o Engine and f a n c o n t r o l and monitoring system o F i r e warning and CO2 e x t i n g u i s h i n g system , Remote c o n t r o l l e d model v a r i a b l e systems o Model p r e s s u r e and temperature i n s t r u m e n t a t i o n system.

A schematic layout of the main tunnel c m t r o l room i s presented i n Figure 4-7. The arrangement of t h e major equipment is shown, i n c l u d i n g t h e t u n n e l c o n t r o l panel, t h e engine and fan c o n t r o l console, t h e on-line computer, p r r s s u r e and temperature r x o r d i n g system, d a t a card punches, computer p r i n t o u t , and TV monitors.

MCOQNNZLL ~ I H C H A F T C-OMPANV 4- I .MDCA4318 FIGURE 4-1 40 x 80 WIND TUNNEL TEST FACILITY Test Section Elevation View

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3c -3" Fi cor I "e!

E lev. 2 - C ' GP76.062: 2 6 2 MCOONNELL AIRCRAFT COMPANY 4-2 MDC 1-131 8 FlCURt 4-3 LARGE S C A t t POWERED MODEL INSTALLED IN 40 FT x 80 F T WiND TUNNEL

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M C O C W U N ~ 1 L R ~ C R I I r r COIUPA N v . .

MDCA4318 FIGURE 4-7 40 x 80 CONTROL ROOM LAYOUT Data Systems Arrangement Test Section Flow Direction Datex I 81 II Control Console Console I SEL 840 Card Computer Reader

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Card Punch I I Datex I I I I Card Punch

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GP76 0622 2 6 7 MCDONNELL AIRCRAFT COMPANY 4-8 blDC A4318 4.2 DATA ACQUISITIOX SYSTEM Two s e p a r a t e Datex Corporation d a t a a c q u i s i t i o n systems are u t i l i z e d i n t h e 40' : i 80' t u n n e l f o r recording tho- balance f o r c e and moment d a t a and t h e model p r e s s u r e and temperature d a t a .

The Datex I system records t h e balance f o r c e d a t a , tunnel test c o n d i t i o n s , and t h e f a n speeds.

The i n d i v i d u a l balance scale readings are measured u t i l i z i n g are measured u s i n g rpm counters.

o p t i c a l scanners and t h e t h r e e f a n speeds These raw d a t a are punched o u t on a n IBN Card Punch l o c a t e d i n t h e on-line computer room a d j a c e n t t o t h e t u n n e l c o n t r o l room. Five readings are taken a t each test p o i n t f o r both t h e balance scale d a t a and f a n speed d a t a . Approximately 10 seconds are r e q u i r e d t o record a given t e s t p o i n t on t h e Datex I system. The Datex I raw d a t a c a r d s are then processed on t h e SEL-840 on-line computer, and t h e r e s u l t s p r i n t e d o u t on t h e t u n n e l c o n t r o l room p r i n t e r .

The Datex I1 system w a s used t o record t h e p r e s s u r e and temperature d a t a .

A l l model p r e s s u r e s were measured u s i n g two Statham Corp. d i f f e r e c t i a l p r e s s u r e transducers mounted i n 48 p o r t , 'ID'' s i z e d s c a n i v a l v e nodules.

The s c a n i v a l v e u n i t s were i n s t a l l e d i n t h e l e f t and r i g h t wing s e c t i o n s of t h e model. The l e f t u n i t contained f i v e ganged modules and t h e r i g h t u n i t contained s i x ganged modules. ii t o t a l of 460 model p r e s s u r e s were measured by :he s c a n i v a l v e systems.

All model teinperature measurements were made u s i n g i r o n / c o n s t a n t a n t'nerzo- couples. The l e a d s of each thermocouple were routed from t h e model through t h e r e f e r e n c e j u n c t i o n box l o c a t e d i n t h e balance room t o a 48 pickup temperature scanner mounted i n t h e Datdx I1 r e c o r d i n g console. The Datex I1 r e c o r d k g system X recorded a l l F X - Y P ~ . ~ ~ and temperature s i g n a l s o n an I B X Card Funch u n i t .

scanning rate o i Q.*3 s?c!:rd p e r measurand was w e d f o r both t h e p r e s s u r e and i,:.!tLhi - D L , r e s u l t i n g i n a r e c o r d i n g time per t e s t p o i n t of temperature dst; approximately 40 ~ ; e = ~ . i j 9 , The Datex I and Datex I1 d a t a were recorded s e p a r a t e l y f o r each t e s t p o i n t , t h e Datex I being recorded f i r s t . The Datex I1 raw d a t a cards were processed on t h e SEL-84G on-line computer.

MCOONNELL AIRCRAFT COMPANY 4- 9 NDC A4318 5 . OUTSIDE STATIC TEST FACILITY The l a r g e s c a l e powered model ground e f f e c t s test program was conducted at t h e NASA-.bes Research Center Outside S t a t i c Test F a c i l i t y , designaced as t e s t s i t e N-249. The outdoor f a c i l i t y is l o c a t e d a t a remote s i t e approximately 1.6 kn number ( 1 mile) from t h e 40' x 80' wind tunnel.

The f a c i l i t y is equipped with a d j u s t a b l e h e i g h t nodel support s t r u t s and provides a smooth ground p l a n e below t h e test model.

The s t r u t arrangement i s similar t o t h a t of t h e 40' x 80' wind tunnel. Height adjv-tment is achieved with interchangeable main s t r u t s e c t i o n s of v a r i o u s l e n g t h s . The t a i l s t r u t is a A remotely d r i v e n t e l e s c o p i n g u n i t t h a t provides a n g l e o f p i t c h v a r i a t i o n .

below-ground s t r u c t u r e supports t h e t h r e e model s t r u t s and t a i l s t r u t d r i v e u n i t .

The f a c i l i t y i n c l u d e s an enclosed trailer t h a t s e r v e s as t h e c o n t r o l room and houses t h e d a t a a c q u i s i t i o n systems. A u x i l i a r y equipment l o c a t e d a t t h e t e s t s i t e includes t h e engine s t a r t e r u n i t , 400 c y c l e X / C power supply, f u e l t a n k e r , and 2n a i r compressor. A plan view s k e t c h o f t h e s t a t i c t e s t f a c i l i t y layout is shown i n Figure 5-1.

5 . 1 TEST AXR.UGEZEXT % d e l Support The model was supported with t h r e e s t r u t s arranged i n t h e sane l c c s t i o n as p r e v i o u s l y used f o r t h e 40' x 80' t u n n e l t e s t s e t u p . A l l nodel and i n s t r u m e n t a t i c n l e a d s wsre a t t a c h e d t o t h e two main support s:ruts and i n s u l a t e d with a s b e s t o s c l o t h wrapping f o r p r o t e c t i o n . S t r u t f a i r i n g s tiere a l s o i n s t a l l e d a t t h e t'no lowest h e i g h t s t e s t e d t o f u r t h e r p r o t e c t t h e model l e a d s from t h e flow f i e l d environment. The nodel was t e s t e d a t t h r e e h e i g h t s above t h e ground p l a n e - 21.0, 8 . 3 , and 3.3 feet. Schematic i l l u s t r a t i o n s o f t h e t h r e e ground h e i g h t arrangenents i n d i c a t i n g t h e major components are shobn i n F i g u r e 5 - 2 . Photos of t h e test model s e t u p s are p r e s e n t e d i n Figures 5-3, 5 - 4 , and 5-5.

Load Cells Three load c e l l s , o n e nsunted on each of t h e support s t r u t s , were used t o neasure t h e f o r c e s on t h e xodel. Each load c e l l was a 3-component s t r a i n p u g e balance, w i t h a 6000 l b normal force,&000 l b a x i a l f o r c e , and 3000 15 s i d e f o r c e c a p a b i l i t y . The load c e l l s were i n s t a l l e d d i r e c t l y below the test c o d e l bet:;een t h e support s t r u t s and nodel mounting ?ads. Yetal shrouds were i n s t a l l e d arzund each load c e l l , and cooling a i r was supplied t o a a i n t a i n near constant l o a d c e l l t e n p e r a t u r e s . The load c e l l and c o o l i n g shrouds can be seen i n t h e p h o t o or' Figure 5-3.

MCDONNELL AIRCRAFT COMPANY 5- 1 MDCA4318 FIGURE 5-1 OUTSIDE STATIC TEST FACILITY PLAN VIEW North Corvus Oil Supply

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\ Ambient Temp, Anemometer, Wind Direction Indicator

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Engine Starter Unit

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Fuel Tanker ?Air Compressor \ /’

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/‘ / Engine Power Supply Unit /’ 4’ iP76 0 6 7 2 d 7 0 MCDONNEII AIRCRAFT COMPANY 5 - 2 MDCA4518 FIGURE 5-2 MODEL GROUND HEIGHT VARIATIONS Outside Static Test Program Cells H = 21.0 Ft GP76 0 8 2 2 278 MCDOUNBLL AIRCRAFT COMPANY 5-3 YDC A4318 =de . t r u u e n t a t i o n

-

The model i n s t r u n e n t a t i o n u t i l i z e d during t h e s t a t i c ground e f f e c t s t e s t ;.ogran was i d e n t i c a l t o t h a t used d u r i n g t h e 40' x 80' wind t u n n e l test.

The q u a n t i t y , type, and l o c a t i o n cf a l l nodel i n s t r u m e n t a t i o n were p r e v i o u s l y d e s c r i b e d ic Section 3 of t h i s r e p o r t . The nodel monitoring and c o n t r o l systems, d e s c r i b e d 4 . 2 , were a l s o t h e same.

i n Section 5 . 2 DATA ACQUISXTICS SYSTEX The measurements made during t h e s t a t i c ground e f f e c t s t e s t p r a g r a n c o n s i s t e d of model p r e s s u r e s , temperatures, fan speeds, and balance f o r c e s . Both d i g i t a l and were u t i l i z e d t o record t h e d a t a .

analog d a t a a c q u i s i t i o n systems The parameters measured and t h e recording s y s t c n s u t i l i z e d are d i s c u s s e d below. X d e t a i l e d sche- ma:ic flow diagram of t h e complete d a t a a c q u i s i t i o n system u t i l i z e d d u r i n g t h e test 9rogram i s presented i n Figdre 5-6.

s t a t i c D i g i t a l Data A c q u i s i t i o n A l l model p r e s s u r e s , temperatures, f a n speeds, and load c e l l cou-.;one~ltf o r c e s were ,.oasured and recorded u t i l i z i n g a Vidar Corporation d i g i t a l d a t a a c q r l s i t i o n system. The nodel p r e s s u r e s were ineasured ir! t h e same manner as during t h e in- tunnel tests u t i l i z i n g d i f f e r e n t i a l 2 r e s s u r e t r a n s d u c e r s i n s t a l l e d T n t x o s e p a r a t e scanivalve s y s i e n s . Teiqc;ature neasurenents were made v i t h iron-constantan 79 t e n p e r a t u r e neasxreme2ts were cor.nec;ed t a :hermoccuples. F o r t p e i g h c of t h e a scanning device f o r s i r g l e c h a n x l multiple:; recording, with t h e o t h e r 31 being recorded on i n d i v i d u a l t e n p e r a t u r e channels. The f a n speeds were measurel 5 t h rpm couniers and recorded 4 3 t i n e s p e r test p o i n t on Vidar using t h r e e scanning channels. Eacn component of t h e t h r e e load c e l l s :.-=IS a l s o recorded 4 8 t h e s ?er test p o i n t usixg t o t a l o f n i c e scanning channels. .Lhe Vidar systen d i g i t a l raw d a t a were recorded 011 a p a p e r tape punch inachice l o c a t e d i n :he on-si:e t r a i l e r .

A schematic layout of t h e d a t a a c q u i s i t i o n system arrangements w i t h i n t h e t r a l l e r u n i t is > r e s e n t e d i n Figure 5-7.

h a l o ? Da:a Acquisition AI1 i n l e t temperat-Jre x e a s u z e m n t s 2nd s e l e c r e d f a n and t i p t c r b i n e e x i t t e a p s r a t u r e s were analog recorded 3n c a g n e t i c tape. I n a d d i t i o n , fa2 s?eed and s e l e c t e d pressures from t h e i n l e t s and f a n ti? t x r b i n e e x i t s were recorded an :;?e tape. Th,se analog ~ e a s u r e m a t s vere nade in o r d e r t o d e t s r x i n e kot17. t h e tise v a r i a n t i n l e t r e i n g e s t i o n c h a r a c t c r i s t i c s and :ne subsequsnt v a r i a t i o n o f t h e propulsion s y s t e a c h a r a c t e r i s t i c s during r e i c g e s t i o n . These z e a s u r s c e n t s vere recorded on two Xmpex CorporP t i o n f / Y tape r e c o r d e r s . I n l e t teiqera:ures were MCDONNELL AIRCRAFT COMPANY 3 - 7 M DC A43 1 8 CI x n n

I

L

T -

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r

c MCDONNELL AJRCFPAFT COMPANY 5-8 MDC A431 8 FIGURE 5-7 OUTSlDE STATE TEST FACILITY CONTROL TRAILER LAYOUT Data Acquisition Systems Arrangement + I

i

I , Table Table

- 14 Track Tape

Paper 1 I Recorder ~ Tape1 - , Punch,

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32 Track ; Tape Recorder

u hrlonitormg

Visicorders VIOAR Data , I System I

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I -

Oscilloscopes 1 Engine '

a Control "nit j bntrpl

I Ambient Cdndition - 4 ! Console I L~~~~ ~ ti i Toward %lode1 M O O N W E L L AlRCUAF T COMPA N V 5 - 9 MDC A4318 recorded on a 32-track % d e l PRLZOO Xnpex t a p e r e c o r d e r , and t h e p r e s s u r e s and f a n exit temperatures were recorded on a 14-track Xodel 1400 Ampex r e c o r d e r . Tempera- t u r e and p r e s s u r e aeasurements were recorded simultaneously on b o t h t h e d i g i t a l Vidar system and t h e analog t a p e r e c o r d e r systems.

S e p a r a t e h i g h response p r e s s u r e t r a n s d u c e r s were used f o r t h e analog p r e s s u r e measureolents. These were connected i n p a r a l l e l with t h e s c a n i v a l v e t r a n s d u c e r s . T i m e code 2nd v o i c e t r a c k s were included on each recorder.

Monitoring Systems Two Honeywell Type 1855 CRT V i s i c o r d e r o s c i l l o g r a p h s were u t i l i z e d t 3 monitor and d i s p l a y s e l e c t e d analog d a t a . One o s c i l l o g r a p h u n i t was used t o monitor d i r e c t l y t h e on-line v a r i a t i o n s i n s e l e c t e d temperatures, p r e s s u r e s , and f a n speed.

The o t h e r u n i t w a s used f o r d i s p l a y i n g playback information recorded on t h e t a p e recorders.

Two 7-bank o s c i l l o s c o p e u n i t s were used t o v i s u a l l y n o n i t o r s e l e c t e d t i m e v a r i a n t s i g n a l s being recorded on t h e two tape recorder u n i t s . Each Sank had t n e c a p a b i l i t y t o d i s p l a y two s i g n a l s , p r o v i d i n g a t o t a l s i g n a l d i s p l a y c a p a b i l i t y of 28 channels.

Ambient a t m s p h e r l c c o n d i t i o n s wer2 measured some d i s t a n c e away f r o = t h e model and displayed on d i g i t a l readout panels l o c a t e d i n t h e t r a i l e r c o n t r o l faom. ATbient temperature, wind v e l o c i t y , vind d i r e c t i o n , and barometric p r e s s u r e verc v i s u a l l y monitored and Sand recorded.

The gas g e n e r a t o r o p e r a t i n g c o n d i t i o n s were v i s u a l l y Eonitored and recarded by hand i n c l u d i n g t n e engine speed, exhaust g a s temperacurs, c i l pressure and temperature, f u e l p r e s s u r e , and engine v i b r a t i o n . The :ur>otik f a n bearing tem- p e r a t u r e s and f a n v i b r a t i o n l e v e l s , t o g e t h e r w i t h f a n rpm (as a backup), were monitored and hand recorded. The above i . s t r u m e n t a t i o n was a l l included on t h e engine and f a n c o n t r o l console in t h e c o n t r o l room. The sane c o n t r o l u n i t w a s used f o r propulsion systen o p e r a t i o n i n 50th t h e 40' x 80' t u n n e l and o u t s i d e s t a t i c t e s t f a c i l i t i e s .

MCOONNELL AIRCRAFT COMPANY 5 - 10 ?LDC A4318 6. i.C?hl> TLXNEL TEST PROGiU?I The test program Conducted i n t h e wind t u n n e l i s d e s c r i b e d i n t h e f o l l o w i n g s e c t i o n s .

6.1 TEST CONDITIOSS Powered l i f t c o n f i g u r a t i o n t e s t i n g was conducted a t s e v e r a l values of t u n n e l dynamic p r e s s u r e a t c o n s t a n t f a n speed t o o b t a i n rha d e s i r e d v a r i a t i o n of j e t v e l o c i t y r a t i o . Aerodynamic l i f t c o n f i g u r a t i o n t e s t i n g was conducted a t a c o n s t a n t v a l u e of dynamic p r e s s u r e over a range of f a n speeds t o o b t a i n a v a r i a z i o n i n i 3 l e t mass flow r a t i o . T e s t dynamic p r e s s u r e v a l u e s , Reynolds number p e r u n i t l e n g t h and nominal j e t e x i t v e l o c i t y r a t i o s and =ass flow r a t i o s are s u m a r i z e d below.

P r e l i m i n a r y T e s t i n g Nominal Jet Reynolds Suaber p e r h i t Dynamic P r e s s u r 2 E x i t V e l o c i t y R a t i o Length a t Standard Amos- (NF/V%= 3600 U Y ) (PSF) (WE?) p h e r i c Conditions P e r Foot Per Yster

1.35 65 0.22 x 106 0.72 x l o 6 0.12

5.5 260 0.43 x l o 6 1 . 4 1 x 106 0.25

12.3 590 0.65 x loc- 3.13 x 106 0.4

21.9 1050 0.87 x 106 2.85 x 1 C 6 0.5 34.2 1640 1.08 x 106 3 . 5 4 x 106 0.6 0.7 49.2 2 360 1.30 x 106 4.27 x 106 Powered L i ' - C o n f i q u r a t i o n Testi3q Xominal Jet Reynoids Sumber per Knit

ACPOS- E x i t - V e l o c i t y R a t i o

Length a t Standard p n e r i c Conditions !NFk'9T, = 35.00 hDY) Per Foot Per Yeter 0.22 x 106 1.35 65 0 . 7 2 x L O 6 0.12 3.34 x 106 1.12 x 106 3.3 160 0.20 7.0 335 0.49 x 106 1.61 x LO6 0.29 1 2 . 3 590 0.65 x 106 2.13 x 10! 0.39 19.2 930 0.81 106 2.66 x loo 0.48 Aerodynamic L i f t Confizura:Lcn Testiag aeynolds :;unber ?er :nit ilynami c P r e s s me Length a t Standard Xt3os- p h e r i c c o n d i t i o n s

- (PSF) (XI n 3

Psr Foot Per Yet?: 34.2 1540 1 "8 x LO5 3 . 5 4 x 106 MCDONNELL AIRCRAFY COMPANY 6-1 Fbdel V a r i a b l e s >!ode1 v a r i a b l e s included twelve c o n f i g u r a t i o n v a r i a b l e s , f a n speeds, a n g l e of a t t a c k and s i d e s l i p .

Preliminary T e s t l n q - I n i t i a l t e s t i n g (Runs 1 through 28) was conducted t o i d e n t i f y b a s i c model c h a r a c t e r i s t i c s ( f l a p e f f e c t c v e n e s s , symmetrical a i l e r o n e f f e c t i v e n e s s and l i f t v a r i a t i o n w i t h angle of a t t a c k ) , t h e induced l i f t charac- terisLics of t h e l i f t c r u i s e u n i t a l o n e , t h e incremental e f f e c t s of t h e nose l i f t u n i t and t h e flow f i e l d c h a r a c t e r i s t i c s a t t h e h o r i z o n t a l t a i l l o c a t i o n . X summary of v a r i a b l e s t e s t e d is shown below. All t e s t i n g was Conducted with the flow survey t a i l o f f .

rake on and t h e h o r i z o n t a l V a r i a b l e Range -4" t o 32" %l a 0" 0" t o 90" 5 LC 0" t o 90" SNL d f 0" t o 30" 0" t o 15" 5a Off, Flow Survey Lake On O'H 0" SR 0 t o 2360 (0 t o A9.2) X/m2 (PSF) q 1600 t o 3600 rpn NF/ 3 Nose Gear Off Nose L i f t Unit I n l e t Covers Open and Covered

Powered L i f t Configuration T e s t i n g - Powered l i f t Configuration f o r c e and

moment d a t a were o b t a i n e d w i t h t h e h o r i z o n t a l t a i l OR and o f f f o r s e l e c t e d combinations of l i f L c r u i s e u n i t geometric defl2c:ion ( 5 ~ ~ 1 and nose l i f t u n i t geometric d e f l e c t i o n (63~) f o r j e t v e l o c i t y r a t i o s ( V 0 / V j ) r e p r e s e n t a t i v e of t h e The e f f e c t of s i d e s l i p angle v a s i n v e s t i g a t e d a t r e s u l t a n t v e c t o r a n g l e ( 3 ~ ) .

angles of a t t a c k of 0' f o r 5Lc/5xL of 90'/90", 5 6 ' / 4 3 " and ? 3 " / 4 3 " and 8" f o r iLc/jyL of 5 6 " / 4 3 " . Horizontal t a i l e f f e c t i v e n e s s was d e t e m i n e d f o r a l l c o n f i g u r a t i o n s t e s t e d . E f f e c t s L f yaw vane d e f l e c t i o n s were t e s t e d a: ~LC/IX,.L = - - * r o n o Y V I Y U ; e f f e c t s of a i l e r o n d e f l e c t i o n and nose 3ear e x t e n s i m xere t e s t z d a t ? L , - / j ~ L = 5 6 " / 4 3 " ; and t h e e f f e c t s of rudder d e f l e c t i o n were t e s t e d a t i L c / < ; j L = F l a p s 3 8 " / 4 3 " . X summary of powered l i f t t o n f i d d r a t i o n ies:ing i s shorn belcw.

and a i l c r o n s were s e t a t 15" and lc)", respec:i*?ely, f o r a i l these t e s t s .

MCDONNELL AIRCRAFT COMPANY 6 - 2 MDC A4318 Nominal V o f V J Range Comments 0.12 t o 0.29 Hor. T a i l o f f ; 3 v a r i a t i o n a t 90 90 Q = 0"; yaw vane e f f e c t i v e r . e s s Very l i m i t e d t e s t i n g 7 1 55 0.12 t o 0.29 Hor. t a i l o f f ; $ v a r i a t i o n a t 56 4 3 0.20 t o 0 . 3 9 a = O", 8 " ; a i l e r o n c o n t r o l power; nose g e a r e f f e c t ; t h r u s t modulation p o r t s open 38 4 3 0.29 t o 0 . 4 8 L i m i t e d t e s t i n g ; rudder E f f e c t i v e n e s s 23 4 3 0.29 t o 0 . 4 8 Hor. t a i l o f f , 3 v a r i a t i o , . a t a = 0" Aerodynarric L i f t C o n f i g u r a t i o n - Aerodynamic l i f t c o n f i g u r a t i o n d a t d were Corrected f a n o b t a i n e d a t a t u n n e l dynamic nressure of 1640 N/n2 ( 3 4 . 2 P S F ) .

sTeeds were v a r l e d from 1600 W>! t o 2700 Z Y t o o b t a i n d e s i r e d mass flow r a t i o v a r i a t i o n . T e s t i n g conducted included h o r i z o n c a l t a i l c o n t r o l power, a i l e r o n c o n t r o l power and rudder c o n t r o l power i n a d d i t i o n t o v a r i a t i o n s i.- a n g l e of a t t a c k and s i d e s l i p .

MCDONNELL AIHCRAFT COMPANY 6- 3 NDC A4318 6 . 2 TEST PROCEDURES S t a t i c C a l i b r a t i o n The propulsion system w a s c a l i b r a t e d a t zero t u n n e l a i r s p e e d w i t h t h e over- head t u n n e l doors open. Each l i f t/cru:se u n i t geometric d e f l e c t i o n was t e s t e d s e p a r a t e l y . S t a t i c f o r c e and uoment d a t a were obtained a t c o r r e c t e d f a n speeds of 2000, 2900, 3600 and b100 R E + ! . I n a d d i t i o n , s t a t i c c a l i b r a t i o n s were repeated f o r ,SLC = 0" a t c o r r e c t e d f a n speeds of 1600, 2130 and 2700 a)!. The nose l i f t u n i t was c a l i b r a t e d f o r nose l i f t u n i t geometric d e f l e c t i o n f r o u 30" t o 104" a t approximately 10" increments a t Z o r r a c t t d f a n speeds of 2000, 2900 and 3600 LEY.

P r e l i a i n a r y Tes t i n q In o r d e r t o provide i n s i g h t i n t o t h e s o u r c e s of power induced f o r c e and moment increments i d e n t i f i e d during t e s t i n g of t h e powered l i f t c o n f i g u r a t i o n , t h e i n i t i a l t e s t i n g examined t h e i n d i v i d u a l e f f e c t s of t h e l i f t / c r L j s e u n i t and nos2 l i f t u n i t s . The l i f t l c r u i s e u n i t s were t e s t e d over a 1arV;e j e t v e l o c i t y r a t i o range w i t h t h e nose lift u n i t i n l e t covered and .jxL = 0" t o o b t a i n t h e e f f e c t of t h e l i f t / c r u i s e u n i t s alone. The nose l i f t u n i t w a s t e s t e d a t a c o r r e c t e d f a n s p e d of 3600 WY f o r S ~ L = 35" t o 90" w i t h t h e l i f t / c r u i s e u n i t a t 5 ~ c = 0" and c o r r e c t e d f a n speed of 3600 X E ? f and w i t h t h e l i f t / c r u i s e u n i t s windmilling (power o f f ) . Also, t h e nose l i f t u n i t was t e s t e d a t 3>iL = 50" w i t h t h e l i f t l c r u i s e u n i t a t SLC = 5 6 " . For t h e s e t e s t s t h e h o r i z o c t a l t a i l =as o f f and t h e flow survey rake was i n s t a l l e d .

Powered L i f t Conf i q u r a t i o n The major test parameters f o r t h e ?owered l i f t c o n f i g u r a t i o n xere j e t v e l a c i t y r a t i o and r e s u l t a n t t h r u s t v e c t o r angle. T e s t i n g was r e s t r i c t o d t o t h e j e t v e l o c i t y r a t i o s which are l i i c e l : ~ t o b e experienced i n f l i g h t a t a r e s u l t a n t t h r u s t v e c t o r angle. I n o r d e r t o provide r e a l i s t i c r e s u l t s t h e c o a b i n a t i o n s of l i f t f c r u i s e u n i t and nose l i f t u n i t georzetric d e f l e c t i o n were selected t o provide an approximate The t a b l e below p i t c h i n g mosent balance a t a p p l i c a b l e j e t v e l o c i t y r a t i o s .

summarizes t h e combinations s e l e c t e d :

- s L c - NL

-

90 90 7 1 55 56 4 3 3d 4 3 2 3 4 3 The o b j e c t i v e s of t h e powered lilt c o n f i g u r a t i a n t e s t i 3 3 werc L O d e t z r a i n e i c l u s z d l i f t and drag c h a r a c t e r i s t i c s , and t o e v a l u a t e 1ongi:udinal ar.d l a t e r a l - d i r s c t i c n a : MCOONNELL APIRCRA F T COMPANY 5-4 YDC A4318 s t a b i l i t y and c o n t r o l c h a r a c t e r i s t i c s .

Aerodynamic L i f t Configuration The primary test parameter f o r t h e aerodynamic l i f t c o n f i g u r a t i o n was t h e mass flow r a t i o (&/A=) f o r t h e l i f t / c r u i s e u n i t f a n i n l e t s . Data were o b t a i n e d a t three values of mass flow r a t i o .

The o b j e c t i v e s o f t h e aerodycamic l i f t configuracion t e s t i n g were t o e s t a b l i s h t h e b a s e l i n e w i t h which t o compare t h e l i f t and drag c h a r a c t e r i s t i c s of t h e powered lift c h a r a c t e r i s t i c s , and t o e v a l u a t e o v e r a l l l o n g i t u d i n a l and 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 and c o n t r o l c h a r a c t e r i s t i c s .

MCDONNELL AJRCRAF T COMPANY 6 - 5 6 . 3 DATA EDUCT= Force and moment d a t a were obtained using t h e s t x d a r d d a t a procedures of t h e 40' x SO' wind tunnel. The measured d a t a (balance d a t a c o r r e c t e d f o r a p p r o p r i a t e weight t a r e s ) were n o t c o r r e c t e d f o r w a l l i n t e r f e r e n c e e f f e c t s , e:qosed t i p s of t h e model support s t r u t s , o r s t r u t i n t e r f e r e n c e e f f e c t s . T h e r e f o r e , t h e measured data i n c l u d e s a l l aerodynanic, propulsion, and i n t e r f e r e n c e f o r c e s a c t i n g on t h e model.

A major o b j e c t i v e of t h e t e s t was t o determine t h e aerodynamic c h a r a c t e r i s t i c s of t h e c o n f i g u r a t i o n ai v a r i o u s t u n n e l and propulsion system o p e r a t i n g c o n d i t i o n s .

I n o r d e r t o i d e n t i f y t h e s e aerodynamic c h a r a c t e r i s t i c s , t h e measured f o r c e s (moments) were s e p a r a t e d i n t o a propulsion sys-en component and an aerodynamic component. The d i r e c t propulsion system conponent is t h e v e c t o r sum of t h e g r o s s t h r u s t and i n l e t ram drag with propulsion u n i t s o p e r a t i n g . The t o t a l aerodynanic component i s t h e d i f f e r e n c e between t h e measured d a t a and t h e d i r e c t propulsion system component. ' h e

aerodynamic component is divided i n t o two components - an aerodynamic component a t a

r e f e r e n c e c o n d i t i o n and a propulsion systcm induced aerodynamic c o q o n e n t .

Gross Thrust and ?lass Flow Xates Xethod of Determining The primary propulsion system periornance parameters t h a t were c a l c u l a t e d during t h e 40' ?I SO' test program included t h e gross t h r u s t ( F G ) . ram d r a g (FxQ~), j e t v e l o c i t y r a t i o ( V o / V j ) , and i n l e t mass flow r a t i o (.%/XHL) f o r a l l t h r e e propul- s i o n u n i t s . - 7 0 b a s i c q u a n t i t i e s were determined from a e a s u r e n e n t s taken 3n t h e f a n and t i p t u r b i n e e x i t r a k e s , namely t h e mass flow r a t e s (72) and i d e a l exi: v e l o c i t i e s ( V I ) f o r each conponen: of each l i f t u n i t . The f a n e x i t and t u r b i n e e x i t mass flows and i d e a l v e l o c i t i e s were c a l t x l a t e d Ltsing a mass-weighting technique f o r each i n d i - v i d u a l rake probe. The e q u a t i o n s f o r c a l c u l a t i x g t-ie propulsion perfornance Faram- eters are as follows:

-!- (6 x v-)

Gross Thrust (FG) = F I x CF = [ ( r i x VI)^^^^

F,LUlCF R a m Drag ( F w i )

= (;TURFj + &&) ?o

J e t V e l c r i t y ( V j ) = FG/ (~TURB + G F ~ ~ ) i n l e t ?lass Flow (Gi) = GTrjRB f wFLL!

- 7 ne mass flow r a t e (G) o f each flat; stream was c a l c u l a t e d f r c n t h e contir,clt]: *qua- t i o n using t h e s t a t i c pressures, t o t a l p r e s s u r e s , aqd t o t a l t e q x r a t u r e s t o compute t h e flow p r a p e r t t e s . The i d e a l e x i t 1:elocit:r ( L ' I ) , nnich i s t h e i s e n t r o p i c o r zaxi- m u m velocit:: t h a t c3n Se obtained expanding t h e e x i t flow through a z e r 9 1 ~ ) ~ s i x z z i e , was a l s o ca1:ula:ed from t h e s t a t i c p r e s s u r e , t o t s 1 pressur?. and t o t a l ter.acr::ures meascrernents used t o complite f l o w p r o p e r t i e s . The t h r u s t c a l i 5 r a t i o n c o e f f i z i c 7 . t (CF) i n t h e above equations WAS tetermined f o r eac3 v e c t c r p o s i r i o n for e a c h a n i t MCDONNELL AIRCRAFT COMPANY 6 - b MDC A4318 from t h e s t a t i c c a l i b r a t i o n s . This parameter was shown t o be e s s e Q t i a l l y c o n s t a n t w i t h fan speed and was assucled t o b e non-varying w i t h forward speed f o r t h e analyses performed in t h i s t e s t program. The t h r u s t c a l i b r a t i o n c o e f f i c i e n t e q u a l s t h e r a t i o of t h e balacce-measured s t a t i c a , d gross tp.rwt (Fg) t o t h e rake- The r e s u l t a n t t h r u s t v e c t o r angles determined measured s t a t i c i a e a l t h r u s t (FI).

from t h e s t a t i c c a l i b r a t i o n s were a l s o shown t o b e e s s e n t i a l l y constant with f a n speed.

Method of Removing Cirect Propulsion Systen Component The propulsion system d i r e c t e f f e c t s were modeled by t h r e e terns, i . e . , t h r u s t v e c t o r a n g l e , g r o s s t h r u s t and mass flow rate. The equations and c o n s t a n t s t h a t were used are ?resented i n Tigure 6-1. I n o r d e r t o use t h e in-tunnel s t a t i c c a l i - b r a t i o n obtained w i t h t h e propulsion u n i t ins:alled in t h e model as t h e b a s i c were r e q u i r e d .

propulsion u n i t performance, t k e following assumptions u The induced loads on :he model when J p e r a t l n g one p r r ? u l s i o n u n i t a t a t i m e are n e g l i g i b l e .

o The fan-rake-neasured mass flow r a t e i s measured a t 2n a c c e p t a b l e l e v e l c f accuracy.

f u n c t i o n D f geometric d e f l e c t i o n angle The r e s u l t a n t t h r u s t a n g l e is a o and is unaffected by tunnel f r e e s t r e a m v e l o c i t y .

Aerodynamic Reference Configuration a s i g n i f i c a n t e f f e c t on d e t e r n i n i n g t h e cagni- The r e f e r e n c e c o n f i g u r a t i o n h a s tude of t h e induced c h a r a c t e r i s t i c s . The r e f e r e n c e c o n f i g u r a t i o n s e l e c t e d f o r t h i s ( 3 ~ c = 0 " , nose f a n i n l e t covered) a t t e s t i s t h e aerodynamic l i f t c o n f i g u r a t i o n u n i t mass flow r a t i o (Xo/Am = 1.0) w i t h a l l aerodgnazic c o n t r o l s d e f l e c t e d t h e same as f o r t h e powered l i f t c o n f i g u r a t i o n . It should b e noted t h a t t h i s d e f i n i t i o n of r e f e r e n c e c o n f i g u r a t i o n i m p l i e s t h a t t h e incremental induced l i f t and drag include t h e e f f e c t s o f t h e e x t e r n a l w e t t e d s u r f a c e of t h e l i f t / c r u i s e n o z z l e s , and t h e nose l i f t f a n i n l s t s and e x i t s as w e l l a s the- induced e f f e c t s of t h e captured stream tubes on the aerodynamic f o r c s s and on t h e o t h e r propuision u n i t s .

A l l powered l i f t c o n f i g u i a t i o n d a t a were obtained w i t h f l a p s l e f l e c t e d 15" and a i l e r m s d e f l e c t e d 10". 3ue t o t e s t l i n i t a t i o n s , t h e aerodynamic L i f t c o n f i g u r a t i o n was t e s t e d :Gith these f l a p and a i l e r o n d e f l e c t i o n s on::: during :he prelinina-: tes:- ing. Consequently, c l s a n c o n f i g u r a t i o n ( f l a p and a i l e r o n a t 0" d e f l e c t i o n ) d a t a was used t o e s t a b l i s h t h e i n c r e n e n t a l e f f e c t s of h o r i z o n t a l t a i l , l a n d i n s g e a r , and f l o w survey rake. A dis-rspancy occurred i n t h e drag d a t a v i t h t h e a d d i t i o n of :he h o r i z o n t a l c a i l a t low angles o f a t t a c k i n t h a t d r a g was reduced i n s t e a d o f i n c r e a s s d MCDONNELL AIRCRAFT COMPANY 6- 7 XDC A4318 as expected ( s e e S e c t i o n 8 . 6 ) . I n e s t a b l i s h i n g t h e d r a g of the r e f e r e n c e configura- t i o n , t h e fol!..?wing test r e s u l t s and assumptions were used: Incremental D l a g C o e f f i c i e n t a t CL = 0 . 2 5 (Based on !u'ing Area) Item

-

0.0100 Nose Gear 0.0070 Flow Survey Rake Horizontal Tail (5H = 0") 0.0000 Flap ( S F = 15") and 0.0165 Aileron ( S a = 1 0 ° / l O o ) Below is t h e summary t a b u i a t i o n of t h e aerodynamic c o e f f i c i e n t s of t h e r e f e r e n c e c o n f i g u r a t i o n a t s e l e c t e d c o n d i t i o n s : o Nose Gear On, 5~ = 0 " , 3F = Eo, ja = 10"/10" Angle of CD CL

Attack (deg) -

0 0.325 0.0705 0.1033 8 0.869 16 1.36 0.246 o H o r i z o n t a l Tail O f f , SF = 15', S a = 1 0 ° / l O o Nose Gear On, Angle of Attack (deg) cL I CD 0 0.355 0.0710 8 0.843 0.1009 16 1.28 0.232 o Nose Gear Off, Flow Su: ey Rake On, i~ = 1 5 ' , ? a = 1 c 1 " Angle o f Attack ( d e % ) 0 0.355 0.0690 0.843 0.0979 1.38 0.229 MCDONNELL A JRCRAF T COMPANY 6- 8 MDCA4318 FIGURE 6-1 EQUATIONS FOR MODELING DIRECT PROPULSION SYSTEM COMPONENTS Force Data Oirect Thrust (No Yaw Vane Deflection) L = FG sin (8 + a ) D = FG COS (e t ai Y - 0 Effect of Inlet Mass Flow Rate L = O D = Vo W COS ($I!¶ .

Y = -Vo W sin ($)/g Moment Data CONSTANTS FOR MODELING DIRECT PROPULSION SYSTEM COMPONENTS FS Model Aefeience Center 276.29 Nose Fan Inlet 114.10 Gas Generator Inlet for Nose Fan 238.10 (kvg Edth Sides) 89.60 -51 24 Lef: Lift Crulse Fan Inlet 282.00 8C.00 -33.00 Gas Generator Inlet for Left Lif: Crbtse F a n 203.90 E9C.J 51 24 Right L,ft Cruise Fan Inlet 282.00 Gas Gemrator Inlet for Right L l f i Cruise Far, 203.90 Nose Fan Thrust Center 11 7.05 Lift Cruise Fan IhrlJst Center 354.50 Left ~ L C = g o 354.50 Rlcjht LC = oc MCDONNELL AIRCRAFT COMPANY 6-9 7 . OUTSIDE STATIC TEST PR3GRM X d e s c r i p t i o n of t h e o u t s i d e s t a t i c test prr-gram, i n c l u d i n g t h e test c o n d i t i o n s , test procedures, and dat.; r e d u c t i o n empioyed, is p r e s e n t e d i n t h e s e c t i o n s t h a t f O l l O i J .

7 . 1 TEST CCSDIFIONS t h e o u t s i d e s t a t i c test progran w a s t o e v a l u a t e t h e ground The purpose of e f f e c t s on t h e t h r e e fan ;/SrOL a i r c r a f t concept. Spt : f i c o b j e c t i v e - of t h e test progran were t o measure t h e e f f e c t s of ground h e i g h t on t o t a l i n r t a l l e d E f t and The model was t e s t e d a t t h r k e grsuad h e i g h t s of 21.0, 8.3, 02 i n l e c reingesLion.

and 3.3 f e e t , correspondlng t o dlD's of 6 . 4 5 , 2.55, and 1.02, r e s p e c t i v e l y .

Landing g e a r h e i g h t f o r t h i s confiqu-ation corresponGs t o a n H / D of approximately 1.02 t h e h e i g h t being aeasured fron rhe i/C n o z z l e exits BP shown i n z i g u r e 5-2.

I n i t i a l t e s t i n g w a s conducted a: t h e 21.0 f o o t r e f e r e n c e n e i g n t , foliovec! by t e s t s a t t h e loves: h i g h t of 3.3 feet, and chen f i n i s h i n g up a t t h e I n t s r m e d i a t e h e i a h t of 8.3 f e e t . The e f f e c t s o f t a n speed v a r i a t i o n s , t h r u s t v e c t o r a c g l e , exhaust j z t s p l a y i n g , and i n l e t s h i e l d i n g were a l s o evaluated. The s p e c i f i c test ConditLcris and model v a r i a 5 l e e v a l u a t e d i n t h e Sround effects test program a r e summarized i n t h e below ta5Ye : 21.0, 8.3, 3.3 o Nodel Height, E {ft) ?Goo, 2900, > b J d . si00

o Fan Speeds, S F / - J T ~ (a!!)

SO, 95, 102 o Sose Y c i t C-eon. D e f l e c t i o n .Angle, ~ N L (deg) o Yozzlil Taw Vane D e f l e c t i o n XnQle, 5y (deg) 0 , 212 a Gas Generator I n l e t S h i e l d S i z e s Large, Snall o I n l e t SLieid Angle, :s (deg) 0 , 4 5 , 90 A11 r e m i n i n g node? t e s t v a r i a b i s s were h e l d a t ;oristzilr values as g i v e n belsw: o L;-ft/Cruise G e o s e t r i c D - f l e c t i o n Angle, ~ L C 90" l o 3 ? A i l s r o n D e f l e c t l a n Xr.g1e, : i 15" 0 F l a p 3e:lecLion Ar.gie. - 1 .

3 Budder D e f l e c t i o n , : R 3"

a . : a r i L m t a l Stabilat,?: Deflectior. :x do

MCDONU6LL ADRCNo F t L'OMPL\NY 7- 1 MDC A4318 I n a d d i t i o n t o t h e ground e f f e c t s t e s t i n g , c a l i b r a t i o n runs tiere made on t h e l i f t / c r u i s e and nose l i f t units for covarison with t h e d a t a o b t a i n e d i n t h e 40' x 80' wind t u n n e l tests. A l s o , a f a n performance map w a s generated f o r t h e T5&8B/S376B p r o p u l s i o n system Tor r e f e r e n c e purposes. It was generated varying t h e f o l l o w i n g parameters using t h e L/H l i f t / c r u i s e u n i t : o P e r c e n t Correctec' Fan Speed, I VF/ v % (RPY) 40 t o 100 o Nozzle E x i t Area, A s ( i n 2 ) 930, 1075, 1385 o Nazzle D e f l e c t i o n Angle, I'LC (deq) 0 The model i n s t r u m e n t a t i o n u t i l i z e d was i d e c t i c a l t o t h a t of t h e 40' x 80' wind t u n n e l tests. The primary q u a n t i t i e s measured d u r i n g t h e s t a t i c r e s t were t h e same as t h o s e considered p r i n a r y i n t h e t u n n e l tests and included t h e following: o Balance-measured Forces and ?loments o Rake-neasured Fan Performance i n c l u d i n g :

-

I d e a l T h r u s t s , F I E - L U and- Fi~t;m.

- Airflow Rates, and GTKRB.

- S o z z l e T o t a l P r e s s u r e Ratio, P R

- Jet V e l o c i t y , L'J

- J e t ' " o t a l Temperature, T , o Rake-uiaasured I n l e t Perforaance i x l u d i n g :

-

i n l e t Xecovery and D i s t o r t i o n , P ~ 2 / ? y , and (PrmX -

- I n l e t T o t a l Temperature Level, T i

I n a d d i t i o n t o the above neasurements, f l o w v i s u a l i z a t i o n runs were a l s o p e r f o m e d a i s e l e c c e d t e s t c o n d i t i o n s . The t e s t procedures u i i l i z e d i n conducting t h e s t a t i c est program a r e d i s c u s s e d i n t h t s e c t i o n t h a t f s l l o w s .

MCDONNELL AIRCRAF f COMPANY 7- 2 HDC A4318 7.2 TEST PROCEDURES The tests conducted during t h i s program i r c l u d e d t h e fol1owi:ig: o I n d i v i d u a l Unit C a l i b r a t i o n Xuns o T5819376B Fan Performacce ?lap Ganeration o Three U n i t Operation Ground E 5 f e c t s T e s t i n g o I n d i v i d u a l Unit O p e r a t i o r Jround E f f e c t ; T e s t i n g o Flow V i s u a l i z a t i o n Tests The s p e c i f i c tests a t t h e 21-foot h e i g h t included t h e following: o Reference Height Ground E f f e c t Tests, a11 u n i t s i n o p e r a t i o n o C a l i b r a t i o n Runs on each u n i t i n d i v i d u a l l y o Fan Yap P e r f o n a n c e Run, l e f t u n i t only A t t h e two lower h e i g h t s t e s t e d , i n d i v i d u a l u n i t performance r u m were conducted on each u n i t s e p a r a t e l y , t h e t h r e e - u n i t o p e r a t i o n ground e f f e c t s performance r u n s , were a c c o q l i s h e d and flow v i s u a l i z a t i o n runs made. A t each ground h e i g h t , a l l model test v a r i a b l e s l i s t e d i n S e c t i o n 7 . 1 were v a r i e d i n a s y s t e m a t i c manner t o provide p a r a m e t r i c d a t a . Ths test sequence and apcroach elnployed f o r each s ? e c i f l c t e s t conducted are describied below.

Fan P e r f o m a n c e >lap A T58/S376B f a n map was g e n e r a t e d a t t h e start of t h e test program u t i l i z i n g The t e s t w a s conducted a: a h e i g h t of 21.0 f e e t h i - h a t h e l e f t l i f t / c r u i s e u n i t .

v e c t c r a n g l e of 0". Three s p e c i f i c n o z z l e areas were i n s t a l l e d on t h e nodel i n t h e c r u i s e =ode and f a n speeds %ere v a r i e d f r o n 1900 t o 4100 r p n ( 4 0 2 t o lOOZ> i n 102 increments. Force and 3o'Jent and s t e a d y s t a t e p r e s s u r e and : e n p e r c i t ~ r e data were recorded.

C a l i b r a t i o n Runs The 'Lndividual u n i t c a l i b r a t i o n runs were F e r f o r a e d a t a h e i g h t of 21.0 Fser (HID = 6 . 4 5 ) following t h e fa? mapping t e s t . The l i f t / c r u i s e u n i t s here v e c t o r e d ~ n d f a n speeds were v a r i e d from 2030 t o 4100 rFn on each. The nGse l i f t u n i t 9Qo . : e 5 tissted a t manually p c s i t i o n e d v e c t o r a n g l e s o f LO, bo, 30, 35 ana 101 d J g r t e s wt..i:= varying tile f a n speed f o r ezcn v e c t o r a c g l e f r o m 2000 co 41uO rpn. S p l a y a n g l e s of 0' and 12" *i2re 31.~0 ' s s : e d For each u n i t u t i l i z i n g t n e yzw vanes. Force and mment b a h z c e d a t a a l m b ... I d i g i x l s t e a d y ;:at2 p r e s s u r e and tsqroeratQre d a t a were r : ? ( . ~ r d e d .

MCDONNt'LL AIRCRAFT COMPANY 7 - 3 ?lDC A4318 Three Unit Ground E f f e c t s T e s t i n q The ground e f f e c t s t e s t sequence w i t h a l l t h r e e u n i t s i n operacion was conducted a t a l l t h r e e ground h e i g h t s . Force and moment d a t a along w i t h i n l e t r e i n g e s t i o n d a t a were recorded € o r each test p o i n t s i n u l t a n e o u s l y , u t i l i z i n g t h e d i g i t a l Vidar and analog t a p e r e c o r d e r s y s t e m . The t e s t sequence v e n t as f o l l o w s : o Beginning z e r o ' s were taken, ambient c o n d i t i o n s recorded, t h e tape r e c o r d e r s and monitoring V i s i c o r d e r s w2re turned on, and t h e t h r e e engines were s t a r t e d and brought t o i d l e . L/C u n i t s were brought up t o t h e i r The two d e s i r e d speeds simultaneously, and then t h e forward u n i t v a s broughc ~p t o speed.

o h%en s t e a d y state c o n d i t i o n s were reached, t h e Vidar d a t a system w a s a c t u a t e d . Ambient c o n d i t i o n s were recorded by hand t h r e e times througnout tne test p o i n t a t 30 second i n t e r v a l s . Appriximateiy 30 seconds werz r e q u i r e d f o r a c q u i s i t i o n of :he Vidar d a t a . X t i m e saquence l o g was recorded by hand f o r t h e tase r e c o r d e r u n i t s t o s e r v e as backtrp t o t h e automatic time code g e n e r a t o r .

The tape r e c o r d e r s and a o n i t o r i n g V i s i c o r d e r were l s f t on during a o complete run, i n c l u d i n g :he e x c u r s i c n s b e w e e n each t e s t p o i n t and d u r i n g t h e engine shutdown sequence. A given i u n l a s t e d anywhere f r o n 5 ninGtes t o 45 a i n u t e s , depending on t h e n d e r of t a s t p o i n t s .

o A t the end of a r u n , t h e engines were brought t o i d l e c o n i i t i 5 n s and t i e t a p e r e c a r d e r s and monitoring Vfsicorder turned o f f . A f t e r a l l o w i n j t h e 2nd z e r c s were : a & : and a n b i e n r f o r engine c o o i i n g , they were s h u t o f f and c o n d i t i o n s recorded.

The only model test v a r i a b l e t h a t was r e o o t e l y c o n t r o l k d xas f a n speea. b given run t h e r e f o r e included only rpm v a r i a t i o n s . The louver a n g l e s , yaw vane s p l a y a n g l e s , and s h i e l d angles were set 3lanually between eacn run.

I n d i v i d u a l Unit Ground E f f e c t s T e s t i n q 73s ground e f f a c t s tests on t h e i n d i v i d u a l uz:-.s iiete c o n d x t e d i n sizilar f a s h i o n a s described above. Each u n i t v2s run i n d i v i d u a l l y over an rps rang2 from ZOCO t o $109 r p , while varyiag t h e louver ar.d/or s p i a y a n 3 l e s . I n addiLion L O :he indi-wyidual it r u x , c o n b i n a t i x rtlcs were ?&rforxed on t h e t i n l i f t . ' : r u i s e u n i t s . The purpose of these tssts was t 5 e v a l u a t e t h e d i f ' L~:sticss jetwee.. :ne i n d i v i d u a l and corsbinad e f f e c t s of m i : o p e r a t i o n on t k s f o r c e and s 0 : 3 e r ~ : oa:a.

I n h t r e i n g e s t i o c d a t a were a l s o atlascred d u r i n s : h s e indi-.tdti;al tests.

MCVONNELL AIRCRAFT COMPO N V 7 - i . W C A4318 Flow V i s u a l i z a t i o n Tests The test model WAS equipped w i t h Corvus o i l i n j e c t i o n n o z z l e s l o c a t e d down- A t strean o f t h e engine i n t h e h o t g a s i n t e r c o n n e c t d u c t i n g on all t h r e e u n i t s .

s e l e c t e d i n t e r v a l s d u r i n g t h e test program, flow v i s u a l i z a t i o n runs were p e r f o . A e d .

The Corvus o i l generated dense smoke i n t h e h o t stream of each U I L ~ , and both normal specd (24 frames/sec) and h i g h speed (250-500 f r a n e s / s e c . ) c o l o r movies were The e n g i n e s were S t i l l c o l o r photos were a l s o taken d u r i n g s e l e c t e d runs.

taken.

t h r o t t l z d t o f a n speeds of 2000 rpm on a l l t h r e e u n i t s , and then t h e Corvus o i l l i n e v a l v e was opened t o supply o i l t o a l l t h r e e u n i t s simultaneously. Run d u r a t i o n w a s approximately 10 secocds each. Eoth t h e n o m 1 and high speed cameras were a c t i v a t e d simultaneously p r i o r t o o i l i n j e c t i o n , and were run through t h e d u r a t i o n of t h e t r a n s i e n t t o stasdy state development of t h e v i s u a l i z e d f l o w f i e l d .

Camera a n g l e s from t h e f r o n t and s i d e of t h e n o d e i were used, b u t n o t s i m u l t a n e o u s l y .

Both t h e V i d a r and analog d a t a s y s t e m were o p e r a t e d j u s t b e f o r e t h e flow v i s u a l i - zacion tests .

MCOONNELL AIRCRAFT cowmatvv 9 - 1 - 3 MDC A4318 9 7.3 DATA REDUCTIOX .

Four s e p a r a t e 4 a t a r e d u c t i o n packages were generated d u r i n g t h e o u t s i d e s t a t i c t e s t program. These d a t a r e d u c t i o n packagcs iccluded t n e following sets of reduced d a t a : o Load Cell and Fan Speed 9 a t z o P r e s s u r e and Temperature 3ata o Propulsf-on System P e r f o m a n c e Data o Analog Temperature and P r e s s u r e Data X d e s c r i p t i o n of each daca package i s presented i n t h e following s e c t i o n s .

- L 1 and Fan Speed Data Package

- - . - i - _ F , . each test p o i n t , t h e load c e l l f o r c e s and far. speeds were recorded tirnes on t h e Vidar system a t a rate of 0.8 second p e r measurement. The a r i t h - A L t j - average of t h e 48 recordings w a s d e t e r a i n e d f o r each component f o r c e and fan speed and p r i n t e d o u t on a summary page t h a t included t h e following reduced d a t a ca:sulatior.s : o S o r n a l , Axial, and Side Forces o L e f t , Sigh:, and Sose Fan Speeds o Pitchin:, Yawing, and Rolling Yocents o Resulc-znt Flow Vector Angles These data were processed on a n IBX 360 computer a t ?i.SSA/;\nes u t i l i t i c g t h e i'idar recorded punched ?aper raw d a t a t a p e s .

-- P r e s s u r e 2nd r e n p e r a t u r e Daca Packase

Fclr each test p o i n t , a l l nodel p r e s s t l r e s and t a p e r s L - 2 s wers each r e s c r d e d one t i m e on t h e Vidar d i g i t a l system using a s c a n r a x of 0 . S second p e r p o r t .

The p r e s s u r e and temperature d a t a package c o n s i s t e d L.f a l i s t i n g of a l l q x a n t i t i e s measured including t h e following: o A l l Yodal P r e s s u r e s i n Ratio Fora (P/Px1;) o All Temperatures i n O F o Lef:, a i g h t , acd Sose Fan S?eeds ( L 3 r e a d i n g s ) Other than t h e r a t i i - i n g of t h e p r e s s u r e n e a s u r s a e n t s , and che conversion of t h e t e u p e r a t u r s 2nd fa- s2eeds t o t h e i r r e s p e c t i v e u n i t s , no f u r t h e r Gaia r e d u c t i o n was Cone. I h f s d a t a package was ais9 ? r o c a s s e J on :;?e 363 ccmputer u s i c g :ne YiCar recorded punched paper caw i a t s tapes.

- Propulsion S:mtex P e r i d n a n c s . Dat? ?ackagc

The propulsion bystem perfarmmc& d a t a ?ackage used i n :he o u t s i d e s:a:i, t e s t program was i d e c t i c a l t o tha: u t i l i z e d ir. t h e SO' -c 80' t u n n s i t e s t ? r e g r a n . Tke MCDONNSLL AIRCRAFT COMPANV 7 - 5 MDC A4318 performance program u t i l i z e d t h e Vidar d i g i t a l o u t p u t from t h e p r e s s u r e and t e n p e r a t u r e d a t a package d e s c r i b e d i n t h e previous s e c t i o n , and c a l c u l a t e d a l l s t e a d y s t a t e p r o p u l s i o n systern perfornance parameters r e q u i r e d f o r a n a l y s i s . The p r i m a r y perfornance parameters t h a t we.-e c a l c u l a t e d and p r i n t e d o u t f o r each l i f t u n i t included t h e following: o Fan and T i p Turbine Performance Data

- T o t a l Corrected Airflows

- Xozzle T o t a l P r e s s u r e R a t i o

- I d e a l Gross Thrust

- Corrected Fan Speeds

- I d e a l Fan Horsepowers

- Sozzle Jet V e l o c i t i e s

- Xozzle Exhaust J e t Temperatures

o I n l e t Pe-formance Data

- Total Zressure Xecovery and D i s c a r t i o n

- Average I n l e t T o t a l T s a p e r a t u r e s and D i s t o r t i o n

were a l s o c a l c u l a t e d and p r i n t e d f o r each component Several a d d i t i o r - a 1 q u a n t i t i e s of aach propu1s;on system u n i t , n a i n l y t h e b a s i c flow p r o p e r c i s s c d c u l a t e d f o r determining t h e above 2arazie:ers. The propulsior. systern d a t a package as also processed or7 :he SASAlAmes ISZ! 360 computer.

Xnaloq P r e s s u r e acd T e q e r a t u r e 3aca The time v a r i a n t pressure and temperature d a t a recorded on t h e two t a p e r e c o r d e r s w a s reduced s i n p i y by p l a y h g back t h e tapes f o r s p e c i f i z rms and r e c o r d i n g t h e analog o u t p u t on o s c i l l o g r a p h pa?er. A11 i n l e t cemperatures and s e l e c t e d f a n and t u r b i n e e s i t p r e s s u r e and temperatures were recorded. The r e d u c t i o n of s e l e c t e d d a t a from s p e c i f i c runs iias actcmplished a t c e r t a i n i n t e r v a l s throughout t h e test program t o ins.:re proper o p e r a t i o n of t h e d a t a a c q u i s i t i o n process.

WCDONNELL AIRCRAFT COMPANY 7 - i 8. WIND TLTNEL TEST RESULTS The r e s u l t s of t h e 40' x 80' Wind Tunnel Test program are p r e s e n t e d i n t h i s s e c t i o a The b a s i c aerodynamic c h a r a c t e r i s t i c s of t h e model, along with t h e p r o p u l s i o n system performance c h a r a c t e r i s t i c s , are p r e s e n t s 2 f o r both t h e powered l i f t and aerodynamic l i f e c o n f i g u r a t i o n s .

The model w a s p r i m a r i l y t e s t e d a t c o n s t a n t c o r r e c t e d f a n speeds, w h i l e v a r y i n g t u n n e l speed t o achieve v a r i a t i o n s i n j e t v e l o c i t y r a t i o .

So a i r f r a n e s t a t i c p r e s s u r e d a t a are presented i n t h i s r e p o r t because of p r o b l e m encountsred i n t h e r e d u c t i o n p r o c e s s .

8.1 PROPULSION SYSTEY STATIC CALIBRATIONS The r e s u l t s of t h e p r o p u l s i o n system s t a t i c c a l i b r a t i o n s are ? r e s e n t e d i n F i g u r e s 8.1-1 through 8.1-13, f o r t h e l e f t and r i g h t l i f t f i r u i s e and n c s e l i f t u n i t s . The balance-measured s t a t i c g r o s s t h r u s t s , t h e rake-measured ideal Qross t h r u s t s , and t h e r e s u l t a n t t h r u s t v e c t o r a n g l e s f o r each & n i t are preser. :d, a l c n g w i t h t h e r e s u l t a n t t h r u s t c a l i b r a t i o n c o e f f i c i e n t s t h a t were detsrmined and u t i - l i z e d throughout t h e forward speed t u n n e l tests. The s t a t i c c a l i 3 r a t i o n s 03 t h e i n d i v i d u a l u n i t s were performed i n t h e 40' x 80' t u n n e l , Y i t h t h e u n i t s i n s t a l l e d i n th.- model.

The balance-measured s t a t i c g r o s s t h r u s t s f d r t h e l e f t , r i g h t , and nose l i f t u n i t s are p r e s e n t e d i n F i g u r e s 8.1-1 through 8.1-4, r e s p e c t i v e l y . The r e s u l t a n t s t a t i c t h r u s t i s p i o t t e d v e r s u s t h e s q u a r e of t h e f a n speed ( t o l i n e a r i z e t h e v a r i a t i o n ) f o r each v e c t o r a n g l a where a f a n speed range was t z s t e d .

The rake-measured i d e a l g r o s s t h r u s t s f o r t h e t h r e e u n i t s are presented i n F i g u r e s 8.1-5 through 8.1-8. These d a t a are p r e s e n t e d i n t h e salne r.anner as t h e f o r c e d a t a , f o r t h e same i d e n t i c a l t e s t p o i n t s and v e c t o r a n g l e s . The rake determined i d e a l g r o s s t h x s t s f o r b o t n t h e f a n and t i p t u r b i n e e x i t fiovs were c a l c u l a t e d u s i n g a n i n d i v i d u a l probe =ass-xeighting technique t l t i l i z i n g t h e fan and t u r b i n e p r e s s u r e and tsrnperature q e a s u r e z e x t s xade a t t h e s t a t o r e s i c s of each.

The tk.rust c a l i b r a t i o n coefficie:l:, d e f i n e d h e r s as t h e r a t i o c f t h e a c t a a l balance-measured s t a t i c t h r u s t t o t h e rake-:.easured i d 5 z l tlzrus: ( F s / F i ) , Yas deternined f o r each geometric v e c t o r a n g l e f o r a l l t h r e e p r s p u l s i c n unizs. A represenLatLve p l o t o f balance t h r u s t ar.d i h a l t h r u s t v e r s u s f a n speed squared is presen:eC i n Tipire 3.1-9 f o r t h e l e f t l i i c , ' c r u i s e c n i t 3: a n o ? z l e d e f i e c t l o n a n g l e or' 90'. The t n r u s t c a l i b r a t i c n coefficient ( C F ) xas di.terained 5 y r z t i o l n g t h e s l c p e s of t h e t x o l i n e a r l y f a i r e d cur..res, as j l l u s t r a t e d i n the fig,ure. The nozzle :hrust c o e f f i c i e n t d e t e r n i n e 3 i n t h i s ;nan;.,er r s 3 a i n s c s x ? tact v i t h c o r r e c t e d MCDONNELL 4IHCRAFT COMPANY 3-1 ?fDC A4318 Comparison o f t h e measured d a t a p o i n t s with t h e i r l i n e a r i z e d f a i r i n g s €an speed.

a l l u n i t s a t each v e c t o r a n g l e over t h e f a n speed range of i n t e r e s t showed a f o r t o l e r a n c e band of + 2 X on t h e t h r u s t c a l i b r a t i o n c o e f f i c i e n t s .

The t h r u s t c a l i b r a t i o n c o e f f i c i e n t s f o r t h e l e f t and r i g h t l i f t l c r u i s e u n i t s As shown i n t h e f i g u r e , both u n i t s have simiiar are p r e s e n t e d i n F i g u r e 8.1-10.

performance c h a r a c t e r i s t i c s w i t h v e c t o r a n g l e , w i t h t h e r i g h t u n i t e x h i b i t i n g approximately 1% g r e a t e r n o z z l e t h r u s t c a l i b r a t i o n c o e f f i c i e n t s a c r o s s t h e v e c t o r e d The nose l i f t u n i t t h r u s t c a l i b r a t i o n r e s u l t s are shown i n F i g u r e 8.1-11.

range.

The n o z z l e v e c t o r i n g performance i s 9 maximum a t t h e 70' v e c t o r a n g l e , as exyectzd.

This is due t o t h e i n s t a l l a t i o n of t h e nose f a n a t a forward t i l t a n g l e of 1 5 " w i t h r e s p e c t t o a w a t e r l i n e .

The r e s u l t a n t t h r u s t v e c t o r a n g l e s p l o t t e d a g a i n s t c o r r e c t e d fan speed f o r t h e l e f t , r i g h t , and :.ose f a n u n i t s are p r e s e n t e d i n F i g u r e s 3.1-12, 8.1-13, and 8.1-14, r e s p e c t i v e l y . As shown i n t h e s e f i g u r e s , t h e r e s u l t a n t t h r u s t v e c t o r a n g l e s f o r each l i f t u n i t vary only s l i g h t l y with f a n s p e e d , and were t h e r e f o r e assumed t o be c o n s t a n t . These r e s u l t a n t t h r u s t v e c t o r a n g l e s were c a l c u l a t e d u s i n g t h e norinal, a x i a l , and s i d e f o r c e components of t h e b a l a n c e d a t a .

The r e s u l t a n t t h r u s t v e c t o r a n g l e s used i3 t h e a n a l y s i s of t h e d a t a i n t h i s r e p o r t are s u m a r i z e d and presented i n F i g u r e s 8.1-15 and 8.1-16, f o r t h e l i f t / c r u i s e u n i t s and nose f a n u n i t s , r e s p e c t i v e l y . As s t a t e d above, t h e flow v e c t o r a n g l e s were assumed t o 3e c o n s t a n t v e r s u s f a n speed f o r each u n i t .

The r i g h t znd l e f t l i f t f c r u i s e v e c t o r a n g l e s were a l n o s t i d e n t i c a l , and t h e r e f o r e t h e sane v a l u e s were u s e d f o r both u n i t s .

The nose u n i t v a l u e s used i n t h e d a t a a n a l y s i s a r e shown i n F i g u r e 8.1-16.

The s t a t i c p i t c h i n g coment v a r i a t i o n f o r b o t h t n e l i f t / c r u i s e and nose l i f t u n i t s as a f u n c t l o n of t h e i r r e s p e c t i v e n o z z l e d e f l e c t i o n a n g l e s i s p r e s e n t e d i n Figures 5.1-17 and 8.1-18.

M C D O W N 6 L . L AlRCRAP T COWPAN Y 3-2 MDCA4318 FIGURE 8.1-1 LEFT LIET/CRUISE UNIT STATIC THRUST Propulsion System Calibration Data v,=o 3 1200

-

e

, j ( l I !

I 1 ! I I / ! i j i l 1 I , 0 2 4 6 8 10 12 14 16 MCDONNHLL A IRCR.4 P T COMPANY 3- 3 MDC A43 1 8

I

FIGURE 8.1-2 RIGHT LIFT/CRUISE UNIT STATIC THRUST Propulsion System Calibration Data V,=O MCDrlNh'Kl-L A fRCRAFT COMPANY 9- !i MDCA4318 FIGURE 8.1-3 NOSE LIFT UNIT STATIC THRUST Propulsion System Calibration Data v,=o E 1200 I 1 I 1 r I ~ I I

i ~ ~

0 2 4 6 8 10 i 2 14 Corrected F a n Soeec Squared. [ N F , ' \ F , 2 . lrprn12 x GP7b 0121 76 0 MCOONN~LL AIRCRAFT CCIMPANY 8- 5 MDC A431 8 I FIGURE 8.1-4 NOSE LIFT UNIT STATIC THRUST Propulsion System Calibration Data V,=Q s 1200 U

e

v) 0 2 4 6 8 10 12 11 e - .

2 6

y 10 . * . irpml

Corrected F i n Speed Squared. [ P . F , \ ; P ? O 0 6 2 2 3 7 MCQONNSLL AIRCIOAPT COmPANV 8- 6 M D C A 4 3 1 8 FIGURE 8.1-5 LEFT LIFTKRUISE UNIT IDEAL THRUST Propulsion System C4ibration Data w,=a U J 1000 soc MCDONNELL AIRCRAF 'I COMPAdY 3- 7 MDC A431 8 FIGURE 8.1-6 RIGHT LIFT/CRUISE UNIT IDEAL THRUST Dropulsion System Calibration Data

I

v,=o

... 111

f lo00

I c 4 6 8 1 0 12 1 4 OO 2 16

Corrected Fan Speed Squared, [ N F / V % ] ~ - (rprn)’ x

GP76.0422 23 MCUONNELL AJRCRAPI C W M N V

-3

MDC A431 8 , - I * FIGURE 8.1-7 NOSE LIFT UNIT IDEAL THRUST Propulsion System Calibration Data V,=3 , I . , ' . ~. ' 1 , , , , ' 0 - 0 2 4 6 8 10 12 14 Corrected Fan Speed Squared. [ N ~ i f i ~ ] - (rpm)* x 10- 6 GP76-0622 20 MCDONNELL AIRCRAFT COMPANY a-9 MDCA4318 FIGURE 8.18 NOSE LIFT UNIT IDEAL THRUST Propulsion System Calibration Data vo=o MCOONNELL AIRCRAFT COMPANY 8- 10 MDC A431 8 FIGURE 8.1-9 THRUST CALIBRATION COEFFICIENT DETERMINATION Left LiftKruise Unit SLC = 900

Corrected Fan Speed Squared, [NF~'\~F,; - ( r m ) 2 x

OP76.0622 301 MCOONNELL. AIRCRAFT COMPANY 8-11 MDC A43 1 8 FIGURE 8.1-10 LIFT/CRUISE UNIT THRUST COEFFICIENTS Propulsion System Calibration Results V , 5 0 '

Geometric Deflection Angle, 6 LC - deg

QP70.0022~~ J MCDONNELL AlcdCRAPT COMPANY

0 - -

MOC A431 8 FIGURE 8.1-1 1 NOSE LIFT UNIT THRUST COEFFICIENTS Propulsion System Calibration Results v,=o Geometric Deflection Angle, hNL e deg GP76 0622 15 &lCDONNhLL AIRCRAFT COMPANY 3-13 MDC A43 1 8 FIGURE 8.1-12 LEFT LIFT/CRUISE UNIT THRUST VECTOR ANGLES Propulsion System Calibration Data v,=o ...

f

2 0 2000 3000 4000

Corrected Fan Speed, N F A , ~ - rpm

GP78 0622 25 MCDONNdtlL AIRCRAFT COMPANY 8- 1 4 MDL A431 8 FIGURE 8.1-13 RIGHT LIFT/CRUISE UNIT THRUST VECTOR ANGLES Propulsion System Calibration Data v,=o $?

u A 3 50 C a

-

5 40

x k L E m c n - 2000 3000 4000

Corrected Fan Speed, NF:~/K - rpm

MCDONNISLL (PIRCIOAFT COMPANY 8- 15 MDCA4318 FIGURE 8.1-14 NOSE LIFT UNIT THRUST VECTOR ANGLES * Propulsion System Calibration Data v,=o o l 2000 ~ 0 0 0 GP76 0622 24 Corrected Fan Speed, NF/d/Bf, . rpm MCDONNELL AIRCRAF t COMPANY

a- 16

MDCA4318 FIGURE 8.1-15 LIFT/CRUISE -_ UNIT THRUST VECTOR ANGLES Propulsion Systsm Calibration Results \

- u,=o

Geometric Deflection Angle, S LC - deg

GP76.0822 14 MCDONNELL AIRCRAFT COMPANY 8-17 MDCA4318 FIGURE 8.1-16 NOSE LIFT UNIT THRUST VECTOR ANGLES Propulsion System Ca I ibration Resu I t s v,=o

Geometric Deflection Angle, S N L - deg

5P76.0622 12 MCDONNELL AIRCmAPY COMPANY 8-15 MDCA4318 FIGURE 8.1-17 STATIC PITCHING MOMENT VARIATION WITH LIFT/CRUISE UNIT GEOMETRIC DEFLECTION ANGLE NF/\/~;-= 3600 RPM LiftiCruise Unit Gecmetric Deflec:;cn, 'SLC . de!

GP76.0622 I66 MCDONNhLL AlRCRAF I COMPANY Q-19 MDCA4318 FIGURE 8.1-18 STATIC PITCHING MOMENIT'VARIATION WITH NOSE i t FT UNIT GEOMETRIC DEFLECTION ANGLE N F / \ ~ = 3600 RPM PICOONNELL AlRCRAF 7 COMPANI' 8-20 M ) ( : A4319 *J 8.2 PROPULSION SPSTOi PEWORw?CE f, The primary performance parameters measured on a l l t h r e e l i f t u n i t s f o r a l l test p o i n t s c o n s i s t e d of t h e gross t h m s t , r a m drag, j e t v e l o c i t y , and mass flow r a t i o . The t y p i c a l v a r i a t i o n of t h e s e parameters w i t h forward speei 's shown i n Figures 8.2-1 and 8.2-2 f o r t h e l e f t l i f t / c r u i s e and nose l i f t u n i t s , r e s p e c t i v e l y .

The d a t a are presen:ed f o r a constant c o r r e c t e d €an speed at an a n g l e of s t t a c k of zero f o r a f i x e d n o z z l e d e f l e c t i o n angle.

The e f f e c t s of a n g l e of a t t a c k on t h e v a r i a t i o n of :he 3 r o p u l s i o n system performance parameters are p r e s e n t e d i n Figures 8.2-3 through 8.2-5. The powered l i f t mode angle of a t t a c k range t e s t e d v a r i e d f r o 3 -4" t o +20°. Per- formance d a t a v a r i a t i o n f o r t h i s a range is p r e s e n t e d i n Figures 8.2-3 and 8.2-4 f o r t h e l i f t / c r u i s e and nose f a n units, r e s p e c t i v e l y , a t t v o d i f f e r e n ; t u n n e l v e l o c i t i e s . As is shown in t h e f i g u r e s , t h e primary performance parameters were C r u i s e mode constant over t h e range of pcwered l i f t a n g l e of a t t a c k t e s t e d .

t e s t i n g was conducted over a Large? angle of a t t a c k range up t o and i n c l u d i n g 32.5'. i i g u r e 8.2-5 shows t h e v a r i a t i o n i n performance a t t h e h i g h e r a n g l e of a t t a c k f o r t h e l e f t l i f t / c r u i s e u n i t , o p e r a t i n g a t a c o n s t a n t tunnel velocLty at t h e t h r e e fan speeds t e s t e d i n thi.c mode. As shown i n t h e f i g u r e , a l l parameters The d a t a a t t h e h i g h e r tiere naintained a t constant values up t o an c t of 16".

angles of a t t a c k begins t o Z e t e r i o r a t e r a p i d l y due t o t h e f l o v separa:ion i n t h e inboard k-ng panel ahead of t h e l i f : / c r u i s e i n l e t .

The effects of forward speed on t h e l i f t / c r u i s e and nose l i f t unit gross t h r u s t and jet v e l o c i t y r a t i o s a t z l t e r n a t e t h r u s t v e c t o r angles are shown i n Figures 8 . 2 - 6 and 8.2-7. As shown i n both f i g u r e s , t h e forward s p e e d tended t o of t h e l a r g e r d e f l e c t i o n a n g l e c o n f i g u r a t i o n s improve t b e t h r u s t r a t i o v a r i a t i o n Comparisons of rhe (those with t h e lcwest s t a t i c t h r u s t s ) a t .a h i g n e r rate.

l i f t / c r u i s e and nose m i t t h r u s t varia:ion w i t h forward speed are shown i n Figure Only a s l i g h t d i f f e r e n c e i n nose u n i t 8.2-8 f o r nozzle d e f l e c t i o n s of 90".

t h r u s t v a r i a t i o n was experienced w i t h t h i s p r o p u l s i o n systern.

A summary cf t h e i n i e t mas; f l o v r a t i o d a t a a s a ftlnction of t u n n e l and fan 8.2-10, and 8.2-11 f o r t h e l e f t and r i g h t speeds is presented Ln Figures 8.2-9, l i f t i c r u i s e and nose l i f t u n i t s , r e s p e c t i v e l y . X summary c o r r e l a t i o n of t h e j e t v e l o c i t y r a t i o d a t a is a l s o p r e s e n t e d i n l i k e s a a n s r f o r t h e t h r e e propulsior.

u n i t s i n Figures 8.2-12 8.2-13 and 8.2-1&.

) MCDONNhLL AlClCRAFt C O M M N Y 5-21 MDCA4318 FIGURE 8.2-1 LIFT/CRUISE UNIT TYPICAL PERFORMANCE CHARACTERISTICS Left LiWCruise Unit O ~ O O N ~ ~ ~ 3 6 0 0 R P M ~ L C P ~

Tunnel Speod, V, - k t s

MCO-AL A4RCRAPT C O M P A N Y 8-2 2 MDC A431 8 FIGURE 8.2-2 NOSE LIFT UNIT TYPICAL PERFORMANCE CHARACTERlSTlCS o 10' NF/\I- 3600 RPM ~ N L p 43' MCDONNhLL AllPCRAPT COMPANY 8-9 3 MDC A431 8 MCOONNELL AIRCRAFT COMPANY 8-24 FIGURE 8.2-4 ANGLE OF ATTACK EFFECT ON NOSE LIFT UNIT PERFORMANCE N ~ / e = 3 6 0 0 RPM 6 ~ ~ = 4 3 ' -4 0 4 8 12 16 20 24 OP73-0122-30

Angle of Attack, a - deg

MCDONNELL AIRCRAFT COMPANY 8-25 MDC A431 8 FIGURE 8.2-5 EFFECT OF HIGH ANLLES OF ATTACK ON PROPULSION SYSTEM PERFORMANCE Left Lift Cruise Unit J = O o Vo=103Kts

Angle oj-Attack. Q - deg

GP76.0622.300 MCDONNELL AIRCRAFT COMPANY 8-26 MDC A431 8 FIGURE 8.26 FORWARD SPEED EFFECT ON LIFT/CRUISE UNIT PERFORMANCE Left Lift/Cru ise Unit a * 0' NF/<= 3600 RPM 0 20 40 60 80 100 120 140

Tunnel Speed, V , - kts

GP76.0622 2 9 MCOONNhLL AIRCRAFT COMPANY 8-27 MDC A431 8 FIGURE 8.2-7 FORWARD SPEED EFFECTS ON VARIOUS NOSE UNIT VECTOR ANGLES

a - 0' N ~/fi = 3600 RPM

1.3 1.2 v) u .

(3 u- d .- c.

1.1 L u) r t 0.9

Tunnel Speed, (;io - kts

OP70-0022 299 MCOONNBLL AIRCRAFT COMPaN Y 8-28 MDCA4318 FIGURE 8.2-8 COMPARISON OF FORWARD SPEED EFFECTS ON THE LIFT/CRUISE vs NOSE LIFT UNITS a =no N F I ~ F ~ = 3600 R PM 1.3 1.2 \ c3 LL d .- Y m 1 . 1 VI . , 5 I - Wl vl 1 .o 0.9 0 20 40 60 80 100 120 140 Tunnel Speed, Vo . kts OP76.0622 298 MCDORJNaLL AlaCHAFY COMPANY 8-29 MDC A431 8.

FIGURE 8.2-9

I

MASS FLOW VARIATIONS WITH FORWARD SPEED Left LiftKruise Unit N F / J ~ ~ , = 3600 RPM a = 0 ' Tunnel Dynamlc Pressure, q . Fsf OP78-0622.237 MCOONNELL AIRCRA F 7 COMPANY 8-30 _I_ ~ -i--- . .

MDC A431 8 FIGURE 8.2-10 MASS FLOW VARIATIONS WITH FORWARD SPEED Right LiftKruise Unit

N~/,,,BT,= 3600 RPM

a = 00 OP70.0622 296 MCDONUEIL AIRCRAFT COMPANY 8-31 MDC&3 i 8

I

FIGURE 8.2-1 1 MASS FLOW VARIATIONS WITH FORWARD SPEED Nose Lift Unit 01 = 0 ' N F 1 6 = 3600 RPM Tunnel Dvnamtc PFesslrre, q . p j f GP76 0622 2 9 6 MCOONNELL AIRCRAFT COMPANY 8-32 MDC A431 8 FIGURE 8.2;12 EFFECT OF FORWARD SPEED ON JET VELOCITY RATIO Left Li f t/Cr uise Unit N F / ~ ~ = 3600 RPM . .

QP76 O b 2 2 204 MCDONNELL AIRCRAF'T COMPANY 8- 3 3 MDCA4318 F I G U R E 8.2-13 E F F E C T OF FORWARD S P E E D ON JET VELOCITY RATIO Right LiftlCruise Unit

NF/ Je7; = 3600 PPM

0.8 0.7 P.6 L

' 0.3

0.2 0.1 0 - 4 8 12 16 20 24

Tunnel Dynamic Pressure, 1 - P S F

GP75 - e 2 2 Z J J MCOONNELL AIRCRAFT COMPANY 8- 34 MDCA4318 FIGURE 82-14 I EFFECT OF FORWARD SPEED ON JET VELOCITY RATIO I Nose Lift for Unit WF/\'BT, = 3600 RPM I)nONNELL. AIRCRAFT COMPANY 0- 35 MDC A4318

8.3 POGIERED LIFT LONFIGLR-TION - IXIUCED LIFT DRAG CHARACTERISTICS

A primary objective o f t h e l a r g e scale powered model test program vas t o e v a l u a t e t h e power induced l i f t and d r a g c h a r a c t e r i s t i c s of t h e powered l i f t c o n f i g u r a t i o n w i t h t h e h o r i z o n t a l t a i l on and o f f . Data w i t h t h e h o r i z o n t a l tail on were o b t a i n e d f o r f i v e combinations o f nose l i f t and l i f t / c r u i s + unit

vector angles: S~c/km = 23O/43O, 38'/43O, 5 6 ' / 4 3 " , 7lo/5S0, and 90°/900. Data

w i t h t h e h o r i z o n t a l t a i l o f f were o b t a i n e d f o r 5~c/3m = 23'/43O, 56"/63", and 90"/90'. A d d i t i o n a l data were o b t a i n e d f o r s e l e c t e d combinations of nose l i f t and l i f t / c r u i s e unit o p e r a t i o n t o e v a l u a t e t h e e f f e c t o f nose u n i t o p e r a t i o n on induced c h a r a c t e r i s t i c s . The r e s u l t s o f t h e s e t e s c s are d i s c u s s e d i n t h e f o l l o w i n g paragraphs.

Induced L i f t and Drag - H o r i z o n t a l f a i l On

F i g u r e s S.3-1 through 9.3-30 p r e s e n t t h e measured l i f t and d r a g d a t a and t h e c a l c u l a t e d p r o p u l s i o n and aerody2amic c o n t r i b u t i o n s f o r :he f i v e powered l i f t con- t a i l on a t 0".

f i g u r a t i o n s w i t h h o r i z o n t a l It w i l l be noted t h a t t h e aerodynamic c o n t r i b u t i o n i s t h e d i f f e r e n c e between two l a r g e v a l u e s . Thus any e r r o r i n t n e measured f o r c e s o r c a l c u l a t e d p r o p u l s i o n system f o r c e w i l l r e s u l t i n s i g n i f i L a n t e r r o r i n t h e aerodynanic c o n t r i b u t i o n .

levels of l i f t ana d r a g are n o t e d on each f i g u r e . The d i f - The r e f e r e n c e f e r e n c e between t h e r e f e r e n c e l i f t o r d r a g and t h e t o t a l aerodynamic l i f t o r d r a g is t h e aerodynanic component induced by t h e p r o p u l s i o n system. The induced l i f t and drag c h a r a c t e r i s t i c s f o r each c o n f i g u r a t i o n w i t h h o r i z o 2 t a l t a i l on are sum- Ylarized i n F i g u r e s 8.3-31 through 8 . 3 - 3 5 . The induced c h a r a c t e r i s t i c s a t 0" a n g l e of a t t a c k f o r t h e f i v e c o n f i g u r a t i o n s are s u m a r i z e d i n Figure 8 . 3 - 3 6 . These d a t a i l l u s t r a t e t h a t , i n general. p o s i t i v e induced l i f t ztd n e g a t i v e induced d r a g a r e i n d i c a t e d f o r a l l f i v e c o n f i g u r a t i o n s . tiowever, i t should be noted t h a t b o t h t h e accuracy t o k-hich t o t a l f o r c e s are measured and increments depend s t r o n g l y on d i r e c t t h r u s t e f f e c t s are c a l c u l a t e d .

Induced L i f t end Drag - H o r i z o n t a l T a i l Off

F i g u r e s 8 . 3 - 3 ; through 8 . 3 - 5 4 preser.t mezsured and c a l c u l a t e d d a t a f o r t h e The comments f o r t h e t h r e e powered l i f t c o n f i g u r a t i o n s with h o r i z o n t a l t a i l o f f .

The induced l i f t and t a i l an c o n f i g u r a t i o n s a r e a l s o a p p l i c a b l e t o t h e s e d a t a .

drag c h a r a c t e r i s t i c s f o r each c o n f i g u r a t i o n a r e surmaiized i n F i g u r e s 8 . 3 - 5 5 The induced c h a i a c t e r i s t i c s a t 0' a n g l e o f a i t a c k f o r t h e t h r e e through 8 . 3 - 5 7 .

c o n f l g u r d t i o n s a r e sumnarized i n Figure S.3-5'3. Comparison o f t h e s e d a t a wit!l MCOONNELL AtWCRAFt COMPANY 3- 36 MDC 84318 t a i l on data, Figure 8.3-36, demonstrates t h a t removing t h e h o r i z o n t a l t a i l i n c r e a s e s t h e induced l i f t , i.e., t h a t indcced l i f t v a l u e s are more p o s i t i v e . T h i s i n c r e a s e , i n l i f t due t o removing t h e t a i l can b e a t t r i b u t e d t o a p r o p u l s i o n induced donwash t a i l .

f i e l d enveloping t h e h o r i z o n t a l During powered l i f t o p e r a t i o n , t h i s l i f t decrement could b e mininized by changing t h e t a i l incidence.

Power Induced E f f e c t s - Nose Unit C o n t r i b u t i o n

During t h e preliininary t e s t i n g , s e l e c t e d combination> o f l i f t / c r u i s e u n i t and t h e e f f e c t of t h e nose l i f t nose l i f t unit d e f l e c t i o n a n g l e s were tested t o e v a l u a t e T e s t d a t a were o b t a i n e d f o r l i f t / unit on induced l i f t and d r a g c h a r a c t e r i s t i c s .

c r u i s e d e f l e c t i o n a n g l e s of O o , 5 6 " , and 90" w i t h t h e nose l i f t u n i t i n l e t and e x i t covered and f o r nose l i f t u n i t d e f l e c t i o n a n g l e s of 50", 70°, 90" w i t h t h e l i f t !

cruise unit a t 0 ' .

8.3-59 through 8.3-64 p r e s e n t t h e measured l i f t and d r a g d a t a and t h e F i g u r e s f o r t h e nose l i f t u n i t c l o s e d c a l c u l a t e d p r o p u l s i o n and azrodynamic c o n t r i b u t i o n s The induced l i f t and d r a g c h a r a c t e r i s t i c s o b t a i n e d from t h e s e d a t a c o n f i g u r a t i o n s .

are sumamrized i n F i g u r e s 8.3-65 and 8.3-66. These two l i f t / c r u i s e u n i t induced c h a r a c t e r i s t i c s are s u b s t a n t i a l l y l a r g e r than t h r e s f a n c h a r a c t e r i s t i c s o b t a i n e d w i t h comparable l i f t / c r u i s e u n i t n o z z l e s e t t i n g s . This r e s u l t is i l l u s t r a t e d i n Ffgures 8.3-67 through 8.3-70 which corcpare t h e t o t a l aerodynamic l i f t and d r a g of two f a n and t h r e e f a n c o n f i g u r a t i o n s . The a d v e r s e e f f e c t of nose l i f t u n i t opera- t i o n on aerodynamic l i f t i s r e a d i l y apparent.

3.3-71 through 8.3-76 p r e s e n t t h e measured l i f t and d r a g d a t a and t h e Figure* calcula'-.ed p r o p u l s i o n and aerodynamic c o n t r i b u t i o n s o b t a i n e d f o r nose l i f t u n i t d e f l e c t i o n s of 5 G 0 , 7 0 ° , and 9 0 ° , w i t h t h e l i f t / c r u i s e u n i t a t 0". The aerodynamic l i f t and d r a g d a t a o b t a i n e d from t h e s e r e s u l t s are summarized i n F i g u r e s 8.3-77 and 8.3-78. The a d v e r s e e f f e c t of nose u n i t o p e r a t i o n on induced l i f t i s e v i d e n t .

The e f f e c t of nose l i f t u n i t o p e r a t i o n on induced l i f t c h a r a c t e r i s t i c s is E~JIC- marized i n F i g u r e 8.3-79 u h i c h shows t h e nose u n i t induced l i f t , dL/FcNL, v e r s u s I n t h i s f i g u r e t h e j e t v e l o c i t y of t h e me u n i t is used i n j e t v e l o c i t y r a t i o .

computing t h e r a t i o , V/Vj. When SLC = 0 , t h e l i f t decrement due t o forward f a n o p e r a t i o n is reduced r e l a t i v e t o t h a t o b t a i n e d vhen S L C = 56" o r ?O" s i n c e t h e induced l i f t f o r ,SLC = 0 is much smaller.

MCOONNIcll AIRCRAFT GOM#'ANY 8-37 MDCA4318

Dyntmic Pressure. q - psf

GP76~0622 6 9 MCDONNhLL AIRCRAFT COMPANV

s- 38

MDCA4318 FIGURE 8.3-2 t

Dynamic Pressure, q - psf

GP78 0622 70 MCDONNhLl AIRCRAFT COMPANY

a- 39

MDC A431 8 FIGURE 8.3-3 Dynamic Pressure, q . psf WCQONNELL AIRCRAFT COMPANV 8-40 MDC A431 8 FIGURE 8.39 DRAG GDYNAMIC PRESSURE 6H = 0' a = * SLC = 23O ~ N L = 43' o J = 20.1' Graphical Summary of Veasured and Calculated Force Data 6 f = 15O 6, = 10°/lOo Nose Gear On N F / ~ B T , = 3600 RPM Dynamic Pressure. q psf GP76 3 6 1 2 7 2 MCDONN&LLL AIRCRAFT COMPANY 3- 'ii MDC A431 8 FIGURE 8.3-5 LIFT vs DYNAMIC PRESSURE = 00 C Y = 1 6 O 6 ~ c = 2 3 O SNL =43* 0 ~ = 2 0 . 1 O Graphical Summary cf Measured and Calculated Force Data 6f 15' 6 , = 10°/lOo Nose Gear On N F / \ T ~ 3600 RPM MCDONNELI AlRCRAF T COMPANY 8-42 MDC A431 8 FIGURE 8.3-6 DRAG vs DYNAMIC PRESSURE 6H = 00 01 = 16O 6LC = 23' 6NL = 43' 0 J = 20.1' Graphical Summary of Measured and Calculated Force Data 6f = 15' 6 , = 10°/lOo Nose Gear On N ~ / f i = 3600 RPM m '0 n -1 -2 -3 -4 WCDONNELL AIRCRAFT COMPLINY 8 - 4 3 MDCA4318 FIGURE 8.3-7 I LIFT vs DYNAMIC PRESSURE 6H = 0' a = 00 6LC= 38' 6NL = 43' O J = 29.2' Graphical Summary of Measured and Calculated Force Data 6f 5 15' 6 , = 10°/lOo Nose Gear On NF/V% E 3600 RPM 8 12 -. 16 20 24 Dynamic Pressure, q . i j j f t P 7 6 3622 4 5 MCDONNBLI AIRCRAFT COMPANY 8-4 5 hlDCA4318 FIGURE 8.3-8 DRAG vs DYNAMIC PRESSURE 6~ = 0' C Y = 0' 6 ~ c = 38O ~ N L = 4 3 ' 0 J = 29.2' Graphical Summary of Measured and Calculated Force Data E, = 75' 6 , = 10°/lOo Nose Gear On F/v 9T, = 3600 RPM MCWNNELL AIRCRAFT COMPANY 8 - 4 3 MDCA4318 FIGURE 8.3-9 LIFT vs DYNAMIC PRESSURE ~ H = O O a = 8 O S ~ c = 3 8 ~ bpdL =43O 0 ~ = 2 9 . 2 O Graphical Summary of Measured arld Calculated Force Data 6f = 15' 6, = 10°/lOo Nose Gear On NF/- 3600 RPM s c c 'c -I

I

2 1 # / / I I 0 4 8 :2 16 20 24 Dynamic Pressure, q . psf GP7b 0622 43 IWCDONN~~LI AIRCRAFT COMPANY 8-k6 MDCA4318 FIGURE 8.3-10 DRAG vs DYNAMIC PRESSURE E H = O ' Q = 8' 6 ~ ~ ~ 3 8 ~ 6 ~ ~ ~ 4 3 ' Uj=29.2' Graphical Summary of Measured and Calculated Force Dnta

S f = 1 5 O 6 , - 10°/lOo Nose Gear On

a -2 -3 -4 MCDONNlZLL AIRCRAFT COMPAN V 8 - 6 7 MDCA4318 FIGURE 8.3-11 LIFT vs YNAMlC PRESSURE b ~ = o ' ( ~ = 1 6 * 6 ~ ~ ~ 3 8 ' SNL =43O Oj=29.2O Graphical Summary of Measured an3 Calculated Force Data 6f 15' 6 , = 10°/lOo Nose Gear On YF/v% = 3600 RPhl L c .- J MDC A431 8 FIGURE 8.3-12 DRAG us DYNAMIC PRESSURE bH = 00 Q = 16O SLC = 38O 1 5 ~ ~ = 43’ 4 J = 29.2’ Graphical Summary of Measured acd Calculated Force Data 15O T a = lo‘/loo Nose Gear On NF/~’%= 3600 RPM I I I I I I I I I !

I I !

, , I

i

I I i I i !

I -4 I 8 12 16 20 24 Dyriamic Pressure, q ~ psi GP76.9622 48 MDC A431 8 FIGURE 8.3-13 LIFT vs DYNAMIC PRESSURE 6” = 0 ’ C Y = 0 ’ 6Lc = 56’ 6NL = 43’ dJ = 4.5’ Graphical Summary of hleasured and Calcuated Force Data 6f = 15’ 5, = 10°/lOo Nose Gear On NF/\’~T,= 3600 RPM

-

I 7 I

I i i

! i I !

i I I t

! I

i

1 i

I

1 1 I

i

‘ I ! I i

1 1 I 1 1 1 i

0 4 8 12 16 20 24 Dynamic Pressure, q . psi CP78 0622 5 7 MCOONNELL AIRCRAFT COMPANY S-50 MDCA4318 FIGURE .. 8.3-14 _. . .

DRAG vs DYNAMIC PRESSURE 6~ = . - I " L Y = 0 ' SLC = 56' b r q ~ = 4 3 ' $5 = 44.5' Grant *:..I Summary of Measured and Calculated Force Data F ' f 15' 6 , * 10°/lOo Nose Gear On NFI\/BT,= 3600 R P M 0 4 8 12 16 20 24 Dynamic Pressure, q . psf GP7S 7622 5 3 MCOONNELL A X R C R A F ~ COMPANY 8-5 1 MDCA4318 FIGURE 8.3-15 LIFT vs DYNAMIC PRESSURE b ~ = 0 ' u = ~ O 6 ~ c = 5 6 ' 6 ~ ~ = 4 3 O O J =44.5' Graphical Summary of Measured and Calcalated Force Data S f = 15' 6 , = 10°/lOo Nose Gear On NFI\;'%= 3600 RPM I 8 12 16 20 24 GP74 0622 5 9 Dynamic Pressure. q . psf MCDOUtV&LL AIRCRAFT COMPANY 3-j2 MDCA4318 FIGURE 8.3-16 DRAG vs DYNAMIC PRESSURE 6H = Oo a = 8' 6 ~ c = 56' ~ N L = 43' 0 J = 44.5' Graphical Summsry of Measured and Calculated Force Data Sf = 15' 6, = 10°/lOo Nose Gear On N~/vlt)T,= 3600 RPM -1 -2 -3 -4 0 4 8 12 1 6 20 24 CP76 i622 60 Dynamic Pressure, q - psf MCDONNKLL AIRCRAFT COMPANY 9-5 3 MDCA4318 FIGURE 8-3-97 LIFT vs DYNAMIC PRESSURE b H = O 0 c\.=16' S ~ c = 5 6 ' 6NL=43' @J=44.5' Graphical Summary of kleasured and Calculated Force Data S f = 150 5, = 100!10~ Nose Gear On N ~ / \ ' r t T , z 3600 RPM Oycamic Pressure. q - D j f C P 7 6 ,2622 5: MCOONUrLL AIRCRAFI COMPANY 3-5; MDCA4318 FIGURE 8.3-18 DRAG vs DYNAMIC PRESSURE

b" = 00 ci = 1 6 O 6 ~ c = 56' 6 N L = 43' d J = 44.5O

Graphical SGrnrnary of Measured and Calculated Force t)..?~ 5 f = 15' '5, = 10°/lOo Nose Gear On - NF/\ = 3600 RPM -1 -2 -3 -4 MCDONNELL AlRCRAFT COMPANY 9-55 MDC A43 18 FIGURE 8.3-19 LIFT vs DYNAMIC PRESSURE

6H = 00 01 = 00 6 ~ c 71' ~ N L = 5 5 O 0 J = 59.8'

Graphical Summary of Measured and Calculated Force Data 6 f = 15' 6 , = 100/100 Nose Gear On N F / ~ % = 3600 RPM MCOOtWNIELL AlRCRAF T COMPANY 8-56 MDC ~ 4 3 1 a F I G U R E 8.3-20 DRAG vs DYNAMIC PRESSURE bH = 0 ' N = 00 6 ~ c = 71' ~ N L = 5 5 ' Oj = 59.8O Graphical Summary of Veasured and Calculated Force D n 6f=1!j0 6 , = 10°/lOo Nose Gear On NF/\%, = 3600 RPM -1 -2 -3 -4 0 4 8 12 16 20 2 4 Dynamic Pressure. q . p;f OP7F 0 $ 2 2 30 MCUONNECL AIRCRAF f COMPANY 8-57 M D Z A131 8 F I G U R E f .3-21 LIFT vs OYNAltllC PRESSURE 6H = 0 ' C I = a0 SLC = 71' A N L = 5 5 O 'jJ = 59.8O Graphical Summary of Measured and Calculated Force Data MCDONNELL AIRCHAFT COMPANY S- j r 3 MDCA4318 FIGURE 8.3-22 DRAG vs DYNAMIC PRESSURE

SH = 00 a = 8' 6 ~ c = 71° ~ N L = 55' 0 J = 59.8'

Graphical Summary of Measured and Calculated Force Data f i f = 15' 8 5 , = 10°/lOo Nose Gear On N F l V T = 3600 RPM MCDONN6LL AIRCRAFT COMPANY 8-59 MDCA4318 FIGURE 8.3-23 LIFT v s DYNAMIC P R E S S U R E dH = 0 ' Q = 16' ~ L C = 71' ~ N L = 55' 0 J = 59.8' Graphical Summary of Measbred and Calcu!ated Force Data '5f 15' ' 5 , = 10°/lOo Nose Gear On N F l v F , 3600 RPM MCOONNELL AIRCRAFT COMPANY 3-60 MDC A431 8 FIGURE 8.3-24 DRAG vs DYNAMIC P R E S S U R E 6H = 0' c \ ' = 16' ~ L C = 71' ~ N L = 5 5 ' 1 ) J = 59.8' Graphical Summary of Measured and Calculated Force Data df = 15' 6 , = 10°/lOo Nose Gear On N~/v"eiT,= 2600 RPM -1 -2 -3 -4 Dynamic Pressure, q - :\sf GP?6 0 6 2 2 34 MCDONNELL A QlHCRA F T CO.HP4 Al Y 3-61 FADCA4318 F I G U R E 8.3-25 LIFT vs DYNAMIC FRESSURE

= 0 ' a = 0 ' 5 ~ c = 9 0 ' ~ N L = 9 0 ' O J 84.7'

Graphical Summary of Measured and Calculated Force D a t a S f = 15' b , = 10°/lOo Nose Gear On NF/\''%= 3600 RPM Dynamic Presiura, q . p s f C P 7 6 ? ? ' J 6 MCDONNELL AIRCRAFT COMPANY a-5: MDC A431 8 FIGURE 8.3-26 DRAG vs DYNAMIC PRESSURE

dH = 00 ct = oo 6LC = 900 dNL - - 900 g J = ~ 4 . 7 ~

Graphics1 Summar./ of Flssured aa7d Cs!cufatzi Force D j t a .5f 1 5 O 5 , = 10°ilOo Nose Gejr On -- NF.'\ :IT = 3603 RPM MDCA4318 FIGURE 8.3-27 .. .. . _ LIFT vs DYNAMIC PRESSURE 6 H = o o 80 SL;=9Co SNL= 90° !IJ = 84.7' Graphical Sumrnarj of Itlezl;red and Calculated Force Gata ,if = 15' ,)a = 10°/lOo Nose Gear On NF/\ = 3600 RPM MCDONNELL AIRCHA F r COMPANY 9 - i k MDCA4318 FIGURE 8.3-25 DRAG vs DYNAMIC PRESSURE S H = O ' a = 8 ' SLC = 900 'SNL = 9 0 ' tj J = 84.7' Graphical Summary oi h'easum! a d Cali:iIxed Foice Cd;3 , S f = 15' = 10°/fOo Nose Gear On NFl\/!'T,= 3600 RPhl -1 -2 -3 - 4 MDC A431 8 FIGURE 8.3-29 LIFT vs DYNAMIC P R E S S U R E

6H = 0' a = 1 6 O 6 ~ c = 9Go S N L = 9 0 ' t.1 J = 84.7'

Graphical Summary of Measured and Calculated Force Data Jf = 15' 6 , = 10°/lOo Nose Gear On N ~ / v v = 3600 RPM MCDONNELL AIF?CRAFT COMPANY 8 - 6 6 MDCA4318 FIGURE 3.3-30 DRAG v s DYP! \?#I:? P R C ? C ' ' R f I ? H = O O ci = 1 6 O SLC = 900 6NL = 900 Oj = 84.7' Graphical Summary oi Measured m d Calculated Force D z t 3 ~ 5 f = 15* '5, = 10°ilOo Nose Gear 3 n N F / ~ F = 3600 RPM MCDONNELL A IRCHA F T C UZMPA N Y 9-67 MDC A33 18 FIGURE 8.3-31 INDUCED LIFT AND DRAG PARAMETERS vs JET VELOCITY RAT!'; AH = 0' 6Lc = 2 3 ' ,TNL = 4 3 ' '.'J = 20.1' Nose Gear On -. .

0 0.1 0.2 0.3 0.4 0.5 0.6 h t ','.?'z:~;y CJI~O. V , V : G P 7 5 : 6 1 " 1 M C D O N ~ E L L AIRCRAI;" COMPANY ,-63 MDC A43 1 8 FIGURE 3.3-32 INDUCED LIFT AND DRAG PARAMETERS vs JET "ELOCITY RATlO 6~ = 0' ISLC = 3 8 ' ~ N L = 43' d J = 2 9 . 2 ' '5f = 15' b , = 10°:lOo Nose Gear 3 n 0.2 0.1 - 0 . 1 0.1 +---

7------ -

-0.1 0 0.1 0.2 0.2 0 4 0 5 0 6 Jet Veisciiy ~ ? C : . O , V,; \/, r;P:6 ? C . ' . .

MCDONNELL AIRCRAFT COMPANY 5-6 3 hlDCA4318 FIGURE 8.3-33 INDUCED LIFT AND D R A G PARAMETERS Y S JET VELOCITY R A T I O 6~ = 0 ' 6 ~ c = 56" S N L = 43' d J = 4 4 . 5 ' 6f = 1 5 O 6 , = 10°/lOo Nose Gear On MCDONAIELL AIRC-RAFT C O M P ~ N Y 8-7$ MDCA4318 FIGURE 8.3-34 INDUCED LIFT AND D R A S PARAMETERS v s JET VELOCITY RATIO d~ = 0 ' ~ L C = 71' SNL = 5 5 ' ?'J = 5 9 . 8 ' b f 15' 5, = 10°/lOo Nose Gear On MCDONNELL AIRCRAFT COMPANY s-7 1.

MDCA4318 FIGUPE 8.3-35 INDUCED LIFT AND DRAG PARAMETERS v s JET VELOCITY RATIO

ISH = 0 ' b ~ c = 9 0 ' ~ N L = 9 0 ' O J = 84.7'

Sf = 15' = 10°/lOo Nose Gear On MDCA4318 FIGURE 8.3-36 SUMMARY OF PROPULSION INDUCED EFFECTS AT 0 ' ANGLE OF ATTACK THREE FAN OPERATION, HORIZONTAL TAIL ON ; 5 ~ = Oo j f = 1 5 O 5, = 10°/lOo Nose Gzar On hl DC A43 1 8 FIGURE 8.3-37 LIFT vs DYNAMIC PRESSURE, HORIZONTAL TAIL O F F & = 00 SLC = 23O &NL = 43' l;'J = 20.1 Graphicrtl Summary o i Measured and Calculated Force D a t a 5 f = 15' ':a = l o o / l o o Nose Gear On c) ...

.- A MCUOWNELL AIRCNAFT COMPANY Y - i i MDC A43 18 F I G U R E 8.3-38 DRAG vs DYNAMIC PRESSURE, HORIZONTAL T A I L OFF

c t = o o 6 ~ c = 2 3 ' ~ N L = 43' t?J = 20.1'

Graphical Summary of hleasursd ;rnd Crtlcuiated Force Data f i f = 1 5 ' :la = 1 0 ° ! l O ' Nose G e x On I NF/\ ' ? T , 3600 HPM -2 -3 0 4 MCD ONNL' L L A If2 C . Y A F f C 0 M PA Al Y ? - 7 5 MDCA4318 FIGURE 8.3-39 LIFT v s DYNAMIC PRESSURE, HORIZONTAL. T A 5 , 01 = 8' S L C = 2 3 ' ~ N L = 43' f / J = 30. I Graphical Summary of bleasured a n d Calcdldted Force D3;a 6 f 15' ' 5 , = l o o / l o o Nose G 2 a r On N F / \ " T , 3600 R P M ?

-1 -2 -3 M D C A 2 3 1 8 FIGURE 8.3-41 LIFT vs DYNAMIC PEESSURE, HORIZONTAL rAlL OFF c? = 16' 6Lc = 23O ANL = 4 3 O 3J = 2 0 . 1 O Graphical SL;mmry of A'easurec! and C~!c:iistzd Fox. Da:a 5 f = 15' 5 , = loo.'loo Nose Gear On r NF/\ GT, = 3600 R P M Cvnamic Pressiirs, q - ;?sf C P 7 6 ,'ir2? 7 3 MCOOMNEAL AIRCRAFT COMPANY 5-; 8 MDCA.1313 FIGURE 8.342 DRAG vs DYNAMIC PRESSURE, HORIZONTAL TAIL O F F ~1==16' S ~ c = 2 3 ' 6NL=43' 0 J = 2 O a 1 O Graphical Summary of Measured and Czlculated Force Data 6f = 1 5 O 5 , = 10°/lOo Nose Gear On NF/\J%= 2600 RPM MCDOHNL7LLAlRCQAFT COMPANY 3-79 MDC A431 8 Oynamic Pressure, c :f MicDOUNEU AIRCRAFT COMPIINV 8-80 MDC A43 18 FIGURE 8.3-44 DRAG vs DYNAMIC PRESSURE, HORIZONTAL TAIL OFF OL S O 0 ~ L C = 560 ~ N L = 4 3 ' t) J = 44.5' Graphical Summary of Measured and Calculated Force Data Sf = 15O Sa = 10°/lOo Nose Gear On N F / K P 3600 RPM I 2 : E

s

% d -1 -2 -3 I 0 4 8 12 16 20 24

Dynamic Pressure, q - psf

GP76.0622 64 MCDONN6LL AlRCRAPT COMPANV 8-81 MDC A431 8 FIGURE 8.345 LIFT vs DYNAMIC PRESSURE, HORIZONTAL TAIL OFF

a = * bLC = 560 bNL = 430 8 J = 44.5'

Graphical Summary of Measured and Calculated Force Data 6 f ~ l ! P 6 , = 10"/lOo Nose Gear On - Dynamic Pressure, q - psf MCDONNELL AIRCRAFT COMWNV a-a2 MDC A431 8 FIGURE 8346 DRAG us DYNAMIC PRESSURE, HORIZONTAL TAIL OFF a = 8' 6LC = 560 8NL = 4 3 ' 8 J 44.5' Graphical Summary of Measured and Calculated Force Data 6f = IS0 6 , = 10°/lOo Nose Gear On -2 -3 -4 0 4 8 12 16 20 24 Dynamic Pressure, q - psf QP76~0622 96 AIIOCRILFT C O M P I W V a-a3 MDC A431 8 FIGURE 8.347 LIFT V I DYNAMIC PSESSURE~HORIZONTALTAIL OFF

L Y = W 6LC = sso bNL * 4 3 ' 8 J = 44.5'

Graphical Summary of Measured and Calculated Force Data sa = loo/loo Nose Gear On 6f = 15O . _.

Graphical Summary of Measured and Calculated Force Data bt 5 11' 6 , = 10°/lOo Nose Gear On --- N F / v K = 3600 RPM s

Q c

i

-1 -2 -3 -4 0 4 a 12 16 20 24 Dynamic Pressure, q psf 0~7e-oe22.84 MDCA4318 FIGURE 8.3-51 LIFT vs DYNAMIC PRESSURE, HORIZONTAL TAIL OFF

at= 80 6LC = 90' 6NL = 90' 8 J 84.7'

Graphical Summary of Measut ed and Calculatsd Force Data

6f 0 15' 6, - loo/loo Nose Gear On

NF/V'% 3600 RPM

,' Reference Llft

0 4 8 12 16 20 24 Dynamic Pressure, q psf OP70 0022 86 klcooiNNIU AQICICRLlFT CoMplONV 8- 88 MDCA4318 FIGURE 8.3-52 8 12 i e 20 24 0 4

Oyniimic Pressure, q - psf

O P ~ O ~ ~ Z ae M O C A43 1 8 FIGURE 8 . 3 4 5 3 LIFTVS DYNAMIC PRESSURE, HORIZONTAL TAIL ~ F F

OL le0 6 ~ c = 90' SNL 900 OJ = 84.70

Graphical Summary of Measured and Calculated Force Data

6, - loo/loo Nom Gear On 6f - IS0

- - ---__ .. .

N F / - - 3600 RPM

a 4 8 12 '6 20 24

Oyriemic Pressure, q - psf

t 3 m . 0 ~ 1 2 . e7 -U AIIIY1ZIICT C W M N V 8-90 MOCA4318 FIGURE 83-64 DRAG vs DYNAMIC PRESSURE, HORIZONTAL TAIL OFF a = i P 6LC = 900 6NL 9oo 8 J 84.7O Graphical Summary of Measured and Calculated Force Data

L f = 150 tia - l@/lfl Nosh Gear On

NF/v% = 3800 R M 0 4 a 12 16 20 24

Dynemrc Pressure, 3 - ,isf

OP78.0612 80 MOCA4318 FIGURE 8.365 INDUCED LIFT AND DRAGXRAMETERS vt JET VELOCITY RATIO HORIZONTAL TAIL OFF b ~ c = 230 ~ N L = # 8 J = m.Io 6f = lso 6a = l@/l@ Nose Gear On 0.1

5 s

9.1

i 5 -0.1

-0.2 0 0.1 0.2 0.3 0.4 0.5 0.6

's Velociw Ratio - V,/VJ

OP%-6022.Se AIPCIIACT -NV 8-92 MDC A431 8 FIGURE 83-56

INDUCED LIFT AND DRAG PARAMETERS vs JET VELOCITV RATIO

HORIZONTAL TAIL OFF i5~c = 5eO ~ N L = r23O B j = 44.5'

6f = 150 8 , = 1aOnoo Nose Gear On 0.2 0 2

z

E

= o k 4.2 MOCA4318 FIGURE 8.347 INDUCED LIFT AND DRAG PARAMETERS vs JET VELOCITY RATIO

HORIZONTAL TAIL OFF 6Lc = 900 6NL = 900 eJ = 84.70

6f le 6, = lOOrra0 Nosa Gear On 0 3 0.2 4 . 1

on

-0.1

Jet Velocitv Ratio - V,/V J

A W A C T C Q M M N V 8-94 MOC A431 8 FIGURE 8.3-58

SUMMARY OF PROPULSION INDUCED EFFECTS AT 00 ANGLE OF ATTACK

THREE FAN OPERATION, HORIZONTAL TAIL OFF _ _ ---

L f - 150 6 , - lOO/l@ Nota Gear On

_ _ _ - _ - - 0.1 - -0.1.

0 0.1 0.2 0.3 0.4 0.5 0.6 Jet Velocity Ratio. V,NJ W784622.*75 MOCA4318 FIGURE 8 . 3 - 5 9 LIFT vs DYNAMIC PRESSURE, FLOW SURVEY RAKE ON

-_I .-__ -- -- -

- 00 aLC = 00 sNL = 00 Inlets covered eJ = io

Graphical Summary of Measured and Calculated Force Data

= le 6 , = 10°/l@ Noso Gear Off

N ~ ~ = ~ ~ M ) R P M

Dynamic Pressure q - psf

MDCA4318 FIGURE 83-60 DRAG DYNAMIC PRESS~IRE, FLOW SURVEY RAKE ON

<Y = 00 6Lc = 00 SNL = oo Inlets covered eJ = 10

Graphical Summary of Measured and Calculated Force Data bf = IS0 6, = lO%@ Nose Gear Off N F / ~ 0 3800 RPM 10 20 30 40 50 60 0~7s-062; as

Dynamic Pressure, q - psf

,vi 3C A43 1 8 FIGURE 8.381

LIFT vs DYNAMIC PRESS~IRE, FLOW SURVEY RAKE ON

a = 00 6 ~ c = 5sO ~ N L = Oo Inlets covered 6 J = 470 Graphical Summary of Measured and Calculated Force Data

6f 1SO 6, - l0O/lO0 NoohGBplOff

NF/-P 3600 RPM - ~ ~ ~ ~ ~ P T C O M C Y W Y 8-98 MOCA4318 FIGURE 8.3-62 DRAG vs DYNAMIC PRESSURE , FLOW SURVEY RAKE ON a= 00 6 L c = 56p ~ N L = Oo Inlets covered 8 J = 470 Graphical Summary of Measured and Calculated Force Data b f = ISo h a = l @ / l O 0 Nose Gear Off N F I ~ * 3600 RPM 0 10 20 30 40 50 60

Dynamic Pressure, q - psf

om.oezz.m -acr coklcylvv a-99 MOCA4318 FIGURE 8 . 3 - 0 3 - LIFT vs DYNAMIC PRESSURE, FLOW SURVEY RAKE ON - - --

6Lc = 900 '6NL = 00 inlets coveted 'eJ = 840

QL = 0O

-Graphical Summary of Measured and Calculated F o - Sf = ISo 6 , = 10°/l@ Nose Gear Off _ _ - - N F - = 3800 RPM 0 10 20 -3!. . - 40 50 60

Dynamic Pressure, q - psf

0P7&0022.100 MDC A43 1 8 FIGURE 8.3-64

DRAG - -- V I DYNAMIC PRE-RUE, FLOW SURVEY - - RAKE ON

a - UQ bLc = 9aO 8NL = 00 Inlets Covered eJ =-&IO

Graphical Summary of Measured and Calculated Force Data

bt - lSo 6 , - 10°/l@ Nore Gear Off

' N ~ / 5 = 3 6 0 0 R P M -1 -2 -3

Dynamic Pressure, q - psf

0170-0122 99 klcookrmu A-ACT CQMCYNV 8- 101 MDC A431 8 FIGURE 8 . 3 - 6 6 SUMMARY OF PROPULSION INDUCED EFFECTS AT Oo ANGLE OF ATTACK -_-- _ _ TWO FAN OPERATION, FLO-W SURVEY RAKE ON 6 , = 10°/lOo Nose Gear On b f = 15' -

--- --- - -

Nora Unit Inlets Covered

SNL - 00

c .O 1.6 a U Y d a 0.8

- E

0.4 -0.4 ' 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 Jet Velocity Ratio, V,/VJ GP70-002241 MCDONNRLL A-RACT SOMPANV 8-102 MDC A431 8

FIGURE 8.3-

SUMMARY OF PROPULSION INDUCED EFFECTS AT Oo ANGLE OF ATTACK

TWO FAN OPERATION, -- FLOW .-- SURVEY - RAKE ON

6f - 16' 6, - l 0 O / l O 0 NoseCear Off

R- Nom Unit Inlets Covered ~ N L = OO - -_- - -- - - 0 . 3 (3 U .

3 0.2 Q -0.1' -a .2.

0 0.1 0.2 0.3 0.4 0.5 0.0 6.7 - ~. - Jet Velocity Ratio, if&, 0P7c0122.02 M DC A43 1 8 FIGURE 8.367

EFFECT OF NOSE LIFT UNIT ON TOTAL AERODYNAMIC LIFT

SLC"580 a=+- - -- --

8f = le 8 , - lO"/lO" Horizontal Tail Off

B

i

7 0 m.

s v

h

c 8.

a.-'

-

c -2 -3 -4 0 4' 8 . i 2 16 20 24

Dynamic Pressure, q - psf

(IM-Oe22.wJa A-RAPT =-MY

a- 105

cc-, MDC A431 8 ..

P ' 0 12 16 20 24 8 . . . -. - Dynamic Pressure, q T a . .

QP78-0622-164 MDC A431 0

MDC A431 8

- _. - - . - _

Dynamic Pressure, q - psf

MDC A43 1 8 FIGURE 8.3-75 LIFT vs DYNAMIC PRESSURE, FLOW SURVEY RAKE ON a=Oo SLC = 00 6 ~ ~ = 9 0 0 t3j * 23.80

Graphical Summay of Measured and Calculated F o m O a t a

6 f = l e O 6 a - lOo/lOO Nom Gear Off

-L A-WAC1 OOMCIONV 8-112 MDC A431 8 FIGURE 82-76 DRAG - - vs DYNAMIC PRESSURE, FLOW SURVEY RAKE ON a = @ SLC'OO S ~ ~ m 9 0 0 8 J * 2 3 . 8 0 Graphical Summary of Measured and Calculated Force Data

S , = W 6, - 100/100 Ncm Gear Off

N F - * 3600 RPM 0 4 8 12 16 20 24

Dynemic Pressure, q - psf

OP7lb~22.120 MDC A $3 1 8 FIGURE 8.3-77 -- . . - -_ - -_ EFFECT OF NOSE LIFT UNIT ON TOTAL AERODYNAMIC LIFT

!LC?!?L!r!?-

6f - le 6, - 10°/lOo Horizontal Tail Off

M-U A B R A C T OOMPLINV 8-114 MDC A431 8 FIGURE 84-78 EFFECT OF NOSE LIFT UNIT ON TOTAL AERODYNAMIC DRAG SLC'OO a = o o Lf = ISo La = 10°/lOo Horizontal Tail Off

i c

0 4 8 12 16 20 24 OP76-06So- 161 MDC A M 1 8 8 . 4 POWERED LIFT CONFIGURATION PITCHING MOMENT C&&XTERISTI;CS T o t a l a i r c r a f t p i t c h i n g moments f o r V/STOL a i r c r a f t must be balanced f o r a l l f l i g h t c o n d i t i o n s from v e r t i c a l takeoff through aerodynamic l i f t f l i g h t .

The c o n f i g u r a t i o n t e s t e d was configured t o provide moment balance a t zero speed w i t h t h e l i f t / c r u i s e and nose l i f t units d e f l e c t e d at 90' ( v e r t i c a l t a k e o f f ) , a t a l l speeds and a p p r o p r i a t e t h r u s t d e f l e c t i o n s in powerec' S i f t f l i g h t and in aerodynamic l i f t f l i g h t .

There are two primary Ditching moment t r i m d u g d e v i c e s in powered l i f t f l i g h t - t h e h o r i z o n t a l t a i l and nose l i f t unit. T e s t i n g was conducted t o o b t a i n s u f f i c i e n t data t o e v a l u a t e t h e proper d i s t r i b u t i o n of t r i m moments from t h e s e two devices.

During t h e preliminary and powered l i f t t e s t i n g , t h e v a r i a t i o n in t o t a l p i t c h i n g moment w i t h nose unit geometric d e f l e c t i o n was evaluated. These tests, w i t h t h e h o r i z o n t a l tail o f f and t h e f l a p s d e f l e c t e d 15' and t h e a i l e r o n s lo', were conducted over a range of dynamic p r e s s u r e s f o r lift/cruise u n i t d e f l e c t i o n s of tests, Figures 8.4-1 through 8.4-3, O', 5 6 O , and 90'. The r e s u l t s of t h e s e demonstrate t h e c a p a b i l i t y of nose l i f t u n i t t o balance p i t c h i n g moments due t o t h e l i f t l c r u i s e units s d aerodynamlc c h a r a c t e r i s t i c s . These r e s u l t s were u t i l i z e d t o select t h e nose l i f t geometric d e f l e c t i o n s f o r t h e powered l i f t c o n f i g u r a t i o n t e s t i n g . It should be noted tha! t h e nose l i f t geometric d e f l e c t l o n s are n o t n e c e s s a r l l y optimum but are c o n s i d e r d r e p r e s e n t a t i v e .

During t h e powered lift c o n f i @ . . t i o n terrting, t h e f i v e powered l i f t configura-

- a range of dynamic p r e s s u r e w i t h t h e h o r i z o n t a l t a i l on

tions were t e s t e d o and, f o r t h r e e c o n f i g u r a t i o n s , with t h e h o r i z o n t a l t a i l o f f . The r e s u l t s of Lnese tests, Figures 8.4-4 through 8.4-8, i l l u s t r a t e t h e c a p a b i l i t y of t h e h o r i z o n t a l tail t o t r i m any r e s i d u a l moments due t o t h e l i f t / c r u i s e and nose l i f t u n i t s .

FLOWSURVEY -EON - - a=@ 1 0 -2 -6 -0 0 20 40 60 80 100 120

Nw-lift Unit Geometric Deflection. ~ N L - deg

OpTI).06Z2. z u MDC A431 8' 0 20 40 60 80 100 120

Nose Lift Unit Geometric Deflection, ~ N L - deg

0 ~ 5 0 0 2 2 - w a MDC A431 8

n

i 6 c

-2 -4 -6 -8 -10 0 20 40 60 80 100 120 ._ . ._

Nose Lift Unit Geometric Deflection,6 ~ c - dag

OP70-~02-2U MC-LL AIRCRL\CT COW-NV 8- 120 MDCA4318 FIGURE 8.4-4 PITCHING MOMENT us DYNAMIC PRESSURE

u m o o bLc 230 6NL = 430 eJ = 20.1~

Measured Data

6, - 10°/100 Noso Gear On 6f = I s 0

N F - f 3600 RPM -1v 4 12 16 20 24 8 .. _ _ ~

Dynamic Pressure, q - psf

Three powered l i f t c o n f i g u r a t i o n s were t e s t e d a t 0" a n g l e o f a t t a c k w i t h v a r i a t i o n s i n a n g l e of s i d e s l i p frQm -4" t o 12' o r 20". The d e t a i l e d r e s u l t s of t h e s e tests are shown io Figures 8.5-1 through 8.5-16 f o r r e s u l t a n t t h r u s t v e c t o r a n g l e s of 20.1", 44.5' and 84.7'. These d a t a show asymmetries a t 0" a n g l e of s i d e s l i p ( f o r example, F i g u r e 8.5-4) and non-uniform v a r i a t i o n w i t h a n g l e of s i d e s l i p (Figure 8.5-12). The probable cause of t h e s e d i s c r e p a n c i e s is t h e d i f f i c u l t y of p r e c i s e l y s e t t i n g e q u a l lif t / c r u i s e u n i t t h r u s t s and maintaining t h e s e equal v a l u e s f o r the 1 0 t o 20 minutes r e q u i r e d t o complete a g i v e n a n g l e of s i d e s l i p test run. The n e t r e s u l t i n g r o l l i n g and yawing moments a t 0" s i d e - slip are t k a result of adding two l a r g e and balancing components from t h e l e f t and r i g k t l i f t / c r u i s e units.

F i g u r e s 8.5-17 and 8.5-18 p r e s e n t t h e d i r e c t i o n a l s t a b i l i t y d e r i v a t i v e s f o r t h e powered l i f t c o n f i g u r a t i o n i n dimensional and c o e f f i c i e n t form.

These d a t a show t h e dominant e f f e c t s of t h e p r o p u l s i o n s y s t e m . d t h e r e s u l t i n g d i r e c t i o r a l I n s t a b i l i t y at low jet v e l o c i t y r a t i o s . The primary s o u r c e of yaw- i n g m o m e n t due t o t h e p r o p u l s i o n system at low j e t v e l o c i t y r a t i o s ls due t o t h e nose f a n inlet mass flow during s i d e s l i p . A t h i g h e r j e t v e l o c i t y r a t i o s , t h e aerodynamic c o n t r i b u t i o n becomes t h e dominant term. F i g u r e s 8.5-19 through 8.5-20 and 8.5-21 through 8.5-22 p r e s e n t t h e lateral s t a b i l i t y and s i d e f o r c e d e r i v a t i v e s , r e s p e c t i v e l y . These d a t a show less p r o p u l s i o n system e f f e c t s t h a n the d i r e c t i o n a l d a t a . is laterally s t a b l e a t a l l b u t one The c o n f i g u r a t i o n 8.5-6, t h e a b r u p t s t a b l e c o n d i t i o n t e s t e d . For t h i s condit?.on, shown i n Figure 4 and 8 degrees a n g l e of s i d e s l i p demonstrates change i n r o l l i n g moment between t h e d i f f i c u l t y of o b j e c t i v e l y determining r e p r e s e n t a t i v e slopes a t 0 ' a n g l e of s i d e s l i p .

L a t e r a l - D i r e c t i o n a l Control E f f e c t i v e n e s s A i l e r o n c o n t r o l e f f e c t i v e n e s s d a t a are p r e s e n t e d i n Figure 8.5-23 ana 8.5-24 f o r one powered lift c o n f i g u r a t i o n w i t h 85 = 44.5 ( 6 ~ c = 5 6 ' , 5 m = 43') a t 0 ' a n g l e o f a t t a c k . The d a t a show t h e e f f e c t of v a r y i n g t h e l e f t a i l e r o n from 4-25" (TED) t o -25' (TEu) w i t h t h e r i g h t aileron a t +loo (TED). A130 shown a r e d a t a f o r e q u a l d i f f e r e n t i a l a i l e r o n d e f l e c t i o n s of - +25' on Figure 8.5-23 only.

Rudder c o n t r o l e f f e c t i v e n e s s i s shown i n Figure 8 . 5 . 2 5 f o r one powered l i f t c o n f i g u r a t i o n w i t h 8J = 29.2' (6Lc - 3 8 ' , 6~ = 43', 6 , - 23') for two dynamic p r e s s u r e v a l u e s over a l a r g e a n g l e of a t t a c k range.

M C - U AIRCRAFT COMPANY 8-126 MDCA4318 FIGURE 8.5-1 YAWING MOMENT V I ANGLE OF SIDESLIP 6~ ~ 0 0 CY=@ 6 ~ ~ s 23O ~ N L = 4 3 ' 0 ~ ~ 2 0 . 1 ~ q = 124PSF Graphical Summary of Measured and Calculated Force and Moment Data

S, = 16" 4 = I O ~ / I O ~ N - Gear On

- -4 0 4 8 12 16 -- oP7e.oezz.1 I O

Angle of Sideslip, - deg

NlODONNILL A8UCRACT COM-NV 8-127 MDC A43 1 8 FIGURE 8.5-2 ROLLING MOMENT AND SIDE FORCE us ANGLE OF SIDESLIP SH 0 ' Q p 0 ' SLC = 2 3 ' ~ N L = 4 3 O 8J = 20.1' q = 124 PSF Graphical Summary of Measured and Calculated Force and Moment Oata

4 - IS0 6 , - 1 0 ° l l O o Nose Gear On

N F / ~ s 3800 Re #VI -4 0 4 8 12 16

Angle of Sideslip, f - deg

0Plb0822~09 MCDONNILL A8RCRAFt COMRdNV 8-128 MDC A43 1 8 -2000 -3wo MCDONNSLL AIRCUACT CONIP14NV 8-123 hlDCA4318 FIGURE 8.54 Rot' ING MOMENT AND SIDE FORCE vs ANGLE OF SIDESLIP 6 ~ = 0 ' a = b o 6 ~ c " n O 6N~'43' 85'20.1' q119.4PSF Graphic ' Summery of Measured and Calculated Force and Momem Data

S, - 16' 6, = 10°/lOo Nose Gear On

- 1 000

-2000 -3000 P -4 0 4 8 12 10

Angle of Sideslip, 9 - deg

MCDONNCiCL AIRCRAFT C89MPIINV 8-130 MDCA4318 FIGURE 8.5-6 YAWING MOMENT vs ANGLE OF SIDESLIP SH'OO a=o0 6 ~ c m 5 6 0 6 ~ ~ ~ 4 3 0 d j = 44.50 q = 3 . 2 PSF Graphical Summary of Measured and Calculated Force and Moment Oats

4 = 16' 6 , = 1 Oo/l 9' Nose dear On

N~/-=3600 RPM 3 w

s

u r

i o

u l c '5 cp ).

-1000 -2000 -3000 4 0 4 8 12 16

Angle of Sideslip, f l - deg

OC78.0822-1 I 2 MDCA431Q MOCA4318 C E o

i

-2000 -3000 4 4 8 12 16

Angle-of Sideslip, 0 - deg

W7O-OWZ*lM MOCA4318 FIGURE 8.58 ROLLING MOMENT AND SIDE FORCE vs ANGLE OF SIDESLIP 6 H " e & m o o 6 ~ c ' W ' 6 ~ ~ ~ 4 3 0 BJ'44.5' q P 7 . 1 P S F Graphical Summary of M e a s d and Calculated Force and Moment Data

&=Is" $=l@/l@ NomGearOn

0 4 8 12 16 - . .-

Angle of Sideslip, 0 - deg

gm.oW2.18(1 MDONNEU A-RACT COMPANY 8-134 MDCA4318 c

P f o

-1OOO -2000 -3000 MicDoNNLtLL AlQplcRAPT C-NY 8-135 MDCA4318 FIGURE 8.5-10 ROLLING MOMENT AND SlOE FORCE VI ANGLE OF SIDESLIP

6 ~ a O " a = @ i5~c=ssO ~ N L = @ O j p 4 4 . 5 0 q * 122PSF

Graphical Summary of Measured and Calculated Force and Moment Data

q45O a,=1oonoo hlmaGearon

Angle & Sideslip, P - beg

MCDON- ARICRLIFT COMPIINV 8-136 MDCA4318 FIGURE 8.5-1 1 YAWING MOMENT vs ANGLE OF SIDESLIP 6 H - P a-0' 6 L c " 9 @ 6 ~ ~ = 9 @ 85'84.7' q = l.4PSF Graphical Summary of Measured and Calculated Force and Moment Data

4 = 150 6, = 10°/lOo Nora Gear On

4 0 4 8 12 16

Angle of Sideslip, 0 - de9

QP76.0022.176 MDCA4318 FIGURE 8.5-12 ROLLING MOMENT AND SIDE FORCE vs ANGLE OF SIDESLIP 6H = 0 ' Q = 0 ' 6LC = 90' ~ N L = 90' 8 J 84.7' q = 1.4 PSF Graphical Summary of Measured and Calculated Force and Moment Data S, = 15O = 10°/lOo Nose Gear On NF/-- 3600 RPM 0 4 8 12 16 Angle of Sideslip, J ~ deg OP76.0822 977 MCWNNRLL A m m A P T COMPIINV

a. L3a

MDCA4318 FIGURE 8.5-13 YAWING MOMENT vs ANGLE OF SIDESLIP 6H = 0 ' a-0' 6LC E 90' ~ N L = 90' 8J = 84.7' Q = 3 . 3 PSF Graphical Summary of Measured and Calculated Force and Moment Data

4 = le 6 , = 10°/100 Note Gear On

N F / ~ = 3600 RPM 4 0 4 8 12 16 Angle of Sideslip, J . deg O P ~ C O ~ Z Z . i 78 -NSU A#RCUACT CO-NV 8-139 MDCA4318 FIGURE 8.5-14 ROLLING MOMENT AND SIDE FORCE vs ANGLE OF SIDESLIP

6H = 0 ' Cy e o* 6Lc 90' 6NL 90' 8 J 84.7' q = 3.3 PSF

Graphical Summary of Measured and Calculated Force and Moment Data

4 - 15' 6 , = 10°/lOo Nose Gear On

NF/%~ 3600 RPM L. c -1 000 -2000 -400 -800 MCDONNELL AIRCRAFT COMPANY 8- 140 MDCA4318 FIGURE 8.5-15 YAWING MOMENT vs ANGLE OF SIDESLIP 6~ = Oo C Y Oo 6LC = 90' ~ N L = 90' 8 J 84.7' q = 7.2 PSF Graphical Summary of Measured and Calculated Force and Moment Data

4 = 15' 6, - 1 Oo/l 0 '

Nose Gear On N~1-m 3600 RPM -400 -800

- 1 200

-4 0 4 8 12 16

Angle of Sideslip, /.I - deg

OP70.0022~180 IWCDONNE&L AdRCRAFT COMPANY 8-141 MDC A431 8 FIGURE 8.5-16 ROLLING MOMENT AND SIDE FORCE vs ANGLE OF SIDESLIP 6H”oo Cr=oo 6LCpgoo 6 ~ ~ ~ 9 0 ’ 8J’84.7’ q = 7.2PSF Graphical Summary of Measured and Calculated Force and Moment Data

& = 15O 6 , = 10°/lOo Nose Gear On

N F 1 - S 3600 RPM -4 0 4 8 12 16 Angle of Sideslip, 3 . deg opie.cezz l e i MCWNN6LL A M m A C T COMPANY

a- 142

MDC A431 8 FIGURE 8.5-17

THREE FAN CONFIGURATION DIRECIJONAL CHARACTERISTICS

6H"oo a-0'-

--

Summary of Measured and Calculated Moment Data

& - ?e 6, = 10°/lOo Nose Gear On

N F 1 - E 3800 RPM wcmoNmmu AIRCRAFT COMPANY 8-143 MDC A43 1 8 FIGURE 85-18 THREE FAN CONFIGURATION DIRECTIONAL CHARACTERICtlCS 6 H ' O O a=oo Summary of Measured and Calculated Moment Coefficient Data

6,; 1 9 6, - 10°/lOo Nose Gear On

0.01 0 Measured Data

I

9 I

/- 4 __

-

# A # a / / I / / A S t h i i i t y AX- Total PropulsionContribution

,?c r -

- .

--

/- Y I 0.01 0

I

Total Aerodynemic Contribution -0.010 0 0.1 0 . 2 0 . 3 0.4 0.5 0.6 O P I I 0622 2SA

Jet Velocity Ratio - V,/VJ

MGDONNBLL AIRCRAFT COMPANY 8- 144 MDC A431 8 FIGURE 8.5-19- THREE FAN CONFIGURATION LATERAL C14ARACTERISTlCS 6H'oo oLpoo Summary of Measured and Calculated Moment Data

6f IS0 6 , - 10°/lOo Nose Gear On

c MCVONIVBLL AIRCIPACT COMPANV 8 - ' 4 5 FIGURE 85-20 THREE FAN CONFIGURATION LATERAL CHARACTERISTICS 6 H P @ asOo Summary of Measured and Calculated Moment Coefficient Data 6 f ~ l P 6,=1@/l@ NoseGeuOn MDC A431 8 FIGURE 8.5-21 . .- - - THREE FAN CONFIGURATION SlDE FORCE CHARACTERISTICS 6 H = O 0 Summary of Measured and Calculated Moment Data

6f = le 6 , - 10°/l@ Nose Gear On

Jet Velocity R&o. VI'VJ 0~7e.oezz.zsa MCDOWNELL AlRCRAPT COMPANY a 4 4 7 MDC A43 1 8 FIGURE 8.5-22 THREE FAN CONFIGURATION SIDE FORCE CHARACTERISTICS 6H'oo a P O o Summary r.f Veasured and Calculated Force Coefficient Data b f * 15O 6, = 10°/lOo Nose Gear On -. .

0 0.1 0.2 0.3 0.4 0.5 0.6 Jet Velocity Ratio, v,,/vJ OP7S. 0622.267 wcu0mNm.L AIRCRAFT COMPANY

a- 148

MDC A431 8

I

C u MDCA4318 0.02 -0.02 -0.04 -0.00

Left Aileron Deflec.,on, baL - deg

OP76.0622.208 MCDONNELL AIRCRAFT COMPANV 8-150 MDC A431 8 FIGURE 8.5-25 EFFECT OF RUDDER DEFLECTION ON LATERAL-DIRECTIONALCHARACTERISTICS

6 ~ ~ 0 0 6Lc-38O e,=a.P

Measured Data

/3 = 0 ' 6, = 2 3 ' 6f - 15' ga = 10°/lOo Nose Gear On

NF--=~&% RPM

Angle of Attack, (11 - deg

OP76.0622 209 8..6

AERODYNAMIC LIFT CONFIGURATION - LONGITUDIXAL CHARACTERISTICS

The aerodynamic l i f t c o n f i g u r a t i o n is d e f i n e d as any c o n f i g u r a t i o n w i t h a l i f t / c r u i s e u n i t geometric d e f l e c t i o n , ~ L C , of O " , t h e nose f a n s h u t o f f , and t h e nose f a n i n l e t covered. This c o n f i g u r a t i o n is used as t h e r e f e r e n c e i n d e t e r m i n i n g t h e level of t h e power induced e f f e c t s . In o r d e r t o e s t a b l i s h t h i s level, t h e aerodynamic l i f t c o n f i g u r a t i o n d a t a m u s t correspond t o a s t a n d a r d i z e d i n l e t mass flow r a t i o . This c o n d i t i o n w a s s e l e c t e d p r i o r t o t h e test as an i n l e t mass flow r a t i o , &/Am, of 1.0. For this test a lift c r u i s e fan speed of 2700 rpm a t a t u n n e l dynamic p r e s s u r e of 34.2 psf r e s u l t e d in a u n i t y i n l e t mass flow r a t i o .

The d a t a @ r e s e n t e d i n t h i s s e c t i o n were used t o e s t a b l i s h t h e r e f e r e n c e lewels p r e s e n t e d i n S e c t i o n 8.3.

It should be n o t e d t h a t t h i s d e f i n i t i o n of r e f e r e n c e c o n f i g u r a t i o n i m p l i e s t h a t the i n c r e m e n t a l induced l i f t and d r a g i n c l u d e t h e e f f e c t s of t h e external w e t t e d s u r f a c e of t h e l i f t / c r u i s e n o z z l e s , and t h e nose l i f t f a n inlets and exits as w e l l as t h e induced e f f e c t s of t h e c a p t u r e d stream tubes on t h e aerodynamic f o r c e s and on t h e c t h e r p r o p u l s i o n u n i t s .

Basic Data Used t o E s t a b l i s h Reference C o n f i g u r a t i o n C h a r a c t e r i s t i c s The b a s i c d a t a used t o e s t a b l i s h t h e r e f e r e n c e l e v e l s f o r t h e powered l i f t c o n f i g u r a t i o n are p r e s e n t e d in t h i s s e c t i o n i n b o t h dimensional and c o e f f i c i e n t form. The dimensional form has been i n c l u d e d i n o r d e r t o emphasize t h e magnitude of t h e p r o p u l s i o n system c o n t r i b u t i o n t o t h e b a l a n c e measured f o r c e s and moments.

Data are p r e s e n t e d f o r t h e t h r e e c o n f i g u r a t i o n s d e s c r i b e d i n t h e following t a b l e .

Rake C o n f i g u r a t i o n 6a 6f 68

- -

1 0 0 on Off

- 7 0 0 Off Off

3 10 15 Off on

Dimensional Data - The dimensional d a t a are p r e s e n t e d i n Figures 8.6-1 through

8.6-9. The l i f t c o n t r i b u t i o n of t h e p r o p u l s i o n system is simp?.y t h e component of gross t h r u s t v e c t o r normal t o t h e r e l a t i v e wind, F G ~ C sin(8LC + a). The p r o p u l s i o n c o n t r i b u t i o n t o t h e l i f t r e s u l t s i n an i n c r e a s e i n b o t h l i f t curve s l o p e and maxi- It is a maximum of 630 l b a t an angle of a t t a c k of 22". This maximm m u m l i f t .

is a r e s u l t of a r e d u c t i o n i n c a l c u l a t e d f a n t h r u s t a t a n g l e s of a t t a c k above wing s t a l l .

MCDONNBII AlRCnAPt COMPANV

a- 15 2

MDC A4318 The p r o p u l s i o n system drag component i n c l u d e s t h e g r o s s t h r u s t component p a r a l l e l t o t h e r e l a t i v e wind and ram drag. The v a r i a t i o n of t h e t o t a l ram drag and t h e g r o s s t h r u s t drag component w i t h a n g l e of a t t a c k are p r e s e n t e d w i t h t h e measured drag data. The d a t a i l l u s t r a t e t h e r e d u c t i o n i n ram d r a g and t h r u s t a t a n g l e s of a t t a c k above wing stall d i s c u s s e d i n t h e p r e v i o u s paragraph.

The p r o p u l s i o n system c o n t r i b u t i o n t o t h e p i t c h i n g moment c o n s i s t s of t h e same components as t h e drag: g r o s s t h r u s t and ram drag. The t h r u s t c o n t r i b u t e s a c o n s i d e r a b l e nose down p i t c h i n g moment due t o t h e t h r u s t l i n e being l o c a t e d above t h e moment r e f e r e n c e c e n t e r . The ram drag, which is t h e r e s u l t of a i r f l o w i n t o b o t h t h e gas g e n e r a t o r and f a n i n l e t s , c o n t r i b u t e s a nose up p i t c h i n g moment.

The r e s u l t i n g t o t a l p r o p u l s i o n system p i t c h i n g moment is nose down.

C o e f f i c i e n t Data - The c o e f f i c i e n t form of t h e t o t a l aerodynamic f o r c e s and moments is p r e s e n t e d i n F i g u r e s 8.6-10 through 8.6-15. These d a t a are t h e c o e f f i - cient form of t h e aerodynamic f o r c e s and moments remaining a f t e r s u b t r a c t i n g t h e p r o p u l s i o n system components. Two irregularities are p r e s e n t i n t h e s e d a t a : o An a p p a r e n t i n c o r r e c t 2" and 6' a n g l e of a t t a c k s e t t i n g o A h i g h e r than a n t i c i p a t e d l i f t c o e f f i c i e n t a t 0" a n g l e of a t t a c k The former can be e x p l a i n e d i f , f o r some reason, t h e a n g l e of a t t a c k s e t t i n g s of 2" and 6 ' were a c t u a l l y c l o s e r t o 2 . 5 " and 6 . 5 " . The latter is more d i f f i c u l t t o r e s o l v e s i n c e t h e c u r r e n t d a t a are s e l f - c o n s i s t e n t . The l i f t c o e f f i c i e n t a t zero angle of a t t a c k , C b , is 0.230 f o r t h e c l e a n wing c o n f i g u r a t i o n w i t h h o r i z o n t a l Previous MCAIR small scale (4.1%) wind t u n n e l t e s t i n g t a i l on a t 0" d e f l e c t i o n .

of a comparable c o n f i g u r a t i o n (Reference ( 3 ) ) i n d i c a t e d a C t o of 0.100. The d i f - f e r e n c e of 0.130 i n t h e C L ~ could be caused by t h e following: Reynolds number d i f f e r e n c e s and scale e f f e c t s o Model support system i n t e r f e r e n c e e f f e c t s o F a b r i c a t i o n accuracy and f l e x i b i l i t y d i f f e r e n c e s e f f e c t s o The MCAIR small scale t e s t i n g of t h i s c o n f i g u r a t i o n had been performed at a 6 based on wing NAC as compared t o t h e p r e s e n t t e s t i n g Reynolds number of 0.6 x 10 performed a t a Reynolds number of 7.5 x 10 based on wing &IC. However, t h e l i t e r a t u r e does not suggest t h a t C L ~ is a f f e c t e d s i g n i f i c a n t l y by Reynolds number.

As mentioned i n Section 6.3, t h e d a t a r e d u c t i o n performed f o r t h i s test did not To a s c e r t a i n t h e magnitude i n c l u d e any model support system i n t e r f e r e n c e e f f e c t s .

of t h i s i n t e r f e r e n c e e f f e c t , a d d i t i o n a l small scale wind t u n n e l t e s t i n g was These tests u t i l i z e d a s c a l e model of t h e A m e s 40' x 80' performed by YCAIR.

MCOONNELL A~HCRAPT COMPANV 8- 15 3 MDC A4318 model support system and t h e small s c a l e model of t h e a i r c r a f t c o n f i g u r a t i o n . The r e s u l t s of t h e s e tests i n d i c a t e d t h e model support system i n t e r f e r e n c e on l i f t c o e f f i c i e n t t o be o n l y 0.025. The one remaining p o s s i b i l i t y , t h a t of f a b r i c a t i o n d i f f e r e n c e s and f l e x i b i l i t y e f f e c t s between t h e tests, cannot be q u a n t i t a t i v e l y e v a l u a t e d a t t h i s t i m e .

I n a d d i t i o n t o t h e high C L ~ , t h e c l e a n wing c o n f i g u r a t i o n has a C L ~ of 1.280 A comparison of l i f t dath from C o n f i g u r a t i o n s a t t h e stall a n g l e of a t t a c k of 15".

2 and 3 (see page 8-152) i n d i c a t e s t h a t d e f l e c t i n g the f l a p s t o 15" and drooping t h e a i l e r o n s t o 10" r e s u l t s i n A C L ~ and A C L w of 0.100 w h i l e m a i n t a i n i n g t h e 15" stall a n g l e of a t t a c k . F i g u r e 8.6-11 p r e s e n t s t h e e f f e c t of t h e h o r i z o n t a l t a i l on drag. The d a t a i n d i c a t e that i n s t a l l a t i o n of t h e h o r i z o n t a l t a i l reduces the aero- dynamic drag. This r e s u l t is probably n o t v a l i d .

The d i s c r e p a n c y may be a n i n d i c a - t i o n of t h e accuracy a s s o c i a t e d w i t h t h e s e p a r a t i o n of t h e p r o p d s i o n and aerodynamic c o n t r i b u t i o n s t o t h e measured drag.

The t a i l on p i t c h i n g moment d a t a i n d i c a t e t h e c o n f i g u r a t i o n t o have s t a t i c l o n g i t u d i n a l s t a b i l i t y ( p r e s t a l l ) w i t h t h e n e u t r a l p o i n t l o c a t e d a t 42% XAC.

D e f l e c t i n g t h e f l a p s t o 15" and drooping t h e a i l e r o n s 10" on t h e t a i l o f f con- f i g u r a t i o n r e s u l t e d i n a nose down p i t c h i n g moment of ACm = - . 0 4 , w i t h a s l i g h t loss i n s t a t i c l o n g i t u d i n a l s t a b i l i t y . N o d a t a were o b t a i n e d € o r t h e t a i l on aerodynamic l i f t c o n f i g u r a t i o n w i t h f l a p s and a i l e r o n s d e f l e c t e d and t h e h o r i z o n t a l t a i l i n s t a l l e d .

H o r i z o n t a l Tail Control E f f e c t i v e n e s s The e f f e c t of h o r i z o n t a l t a i l d e f l e c t i o n on t h e l o n g i t u d i n a l c h a r a c t e r i s t i c s is p r e s e n t e d i n Figures 8.6-16 through 8.6-18. It is shown t h a t t h e c o n f i g u r a t i o n has s t a t i c l o n g i t u d i n a l s t a b i l i t y (prestall) and t h a t n e g a t i v e (TEU) s t a b i l a t o r d e f l e c t i o n s would be u t i l i z e d f o r trimming. Data f o r t h e p o s i t i v e d e f l e c t i o n s i n d i c a t e a s t a l l e d h o r i z o n t a l tail.

MCOONNELL AIRCRAFT COMPANY 8-154 MDC A431 8 FIGURE 8.6-1- LIFT vs ANGLE OF ATTACK

6~ = 0 ' q = 34.2 PSF d ~ c = 0 ' ~ N L = SEALED 85 = 1'

Graphical Summary of Measured and Calculated Force Data Run 116 6f 0 ' 6a Oo/Oo Nose Gear Off N ~ 1 % 2700 RPM OP76-0822 168 MCDONNPLL AmCRAPT COMPANY 8-155 M DC A43 1 8 FIGURE 8.6-2 DRAG vs ANGLE OF ATTACK 6 ~ m O ' qm34.2PSF 6 ~ c s O O GNL'SEALED 8 j = l 0 Graphical Summary of Measured and Calculated Force Data Run 116 b f - O o 6,-O0/Oo NoseGearOff

&/fi0 5 2700 RPM

MCDONNBIL AJRCRAPT COMPANY 8-156 MDCA4318 FIGURE 8.6-3 PITCHING MOMENT vs ANGLE OF ATTACK 6~ 0' q = 34.2 PSF 6 ~ c 0 ' ~ N L = SEALED 6 j = 1' Graphical Summary of Measured and Calculated Moment Data

Run 116 6f 0' 6 , - Oo/Oo Nose Gear Off

N ~ / f i m 2700 RPM MC-NmLL AIRCUACT COMPANY 8-157 MDC A431 8' FIGURE 8.64- LIFT vs ANGLE OF ATTACK, HORIZONTAL TAIL OFF q 34.2 PSF 6 ~ c = Oo ~ N L SEALED 8 J 1' Graphical Summary of Measured and Calculated Force Data Run 140 6 f = 0 ' 6, = Oo/Oo Nose Gear Off N F / ~ O 2700 RPM

Angle of Attack, a - deg

OP76.0622 160 MDCA4318 FLOUR€ 8.6-5 DRAG vs ANGLE OF ATTACK, HORIZONTAL TAIL OFF

q ~134.2 PSF 6 ~ c Oo ~ N L = SEALED 6 J = 1'

Graphical Summary of Measured and Calculated Force Data

Run 140 df - Oo 6 , - Oo/Oo Nose Gear Off

NF/-p 2700 RPM

F

n -1 -2 -3 R 1 G o o N I v B L L AlRCR O F T COMPLINV 8-159 MOCA4318 FIGURE 3.66 PlTCHlNG MOMENT vs ANGLE OF ATTACK, HORIZONTAL TAIL OFF q 3 4 . 2 PSF 6 ~ c = OO ~ N L = SEALED 0 J = 1' Grphical Summary of Measured and Calculated Moment Data

Run 140 6 f oo 6 a - @/O0 Nose Gear Off

NF%= ZMO RPM M D O N N R U A-mACT COMPANY 8-160 MDCA4318 FIGURE 8.6-7 LIFT vs ANGLE OF ATTACK. FLOW SURVEV RAKE ON q = 34.2 PSF 6 ~ c 00 ~ N L = 00 Inlets C o v d @ J = 1 ' Graphical Summary of Measured and Calculated Force Data

Run 26 S, - 1 9 = 10°.'lOo Nasa Gear Off

NFI- 0 2700 RPM W-NELL. AIRCmAPT COMPANY 8- 16 1 MDCA4318 FIGURE 8.68 DRAG vs ANGLE OF ATTACK, FLOW SURVEY RAKE ON q = 3 4 . 2 PSF b ~ c = Oo ~ N L = 0' Inlets Covered 6~ * lo Graphical Summary of Measured and Calculated Force Data Run 26 = 1 9 6 , = 10°/lOo Nose Gear Off N F / ~ = 2700 RPM MDC A43 I8 FIGURE 8.69

PITCHING MOMENT YS ANGLE% ATTACK, FLOW SURVEY RAKE ON

q = 3 4 . 2 PSF blc Oo ~ N L = 0 ' Inlets Covered 6 J = lo Graphical Summary of Measured and Calculated Moment Data Run26 $=le 6a=100/100 NaeGearOff NF/<= 2700 RPM

Angle of Attack. Q - deg

MDC A431 8 FIGURE 8.6-10 EFFECT OF HORIZONTAL TAIL ON LIFT COEFFICIENT vs ANGLE OF ATTACK Q 3 4 . 2 PSF 6 ~ c * 0 ' SNL =SEALED 8J = 1 ' Direct Thrust Effects Removed 6f = Oo 5, = Oo/Oo Now Gear Off N F - 2700 RPM 1 -6 1.4 1 2 1.0

z

2 0.8

A u 0.6 0.4 0.2 4 0 4 8 12 16 20 24

Angle of Attack, a - deg OP76.0622. 140

WC-NRU A-ACT COMPLINY 8-164 MOCA4318 MCDONN6LL AlRCRAFI COMPANY 8-165 MOCA4318 FIGURE 83-12 EFFECT OF HORl2ONTAL TAIL ON PITCHING MOMENT COEFFICIENT vs ANGLE OF ATTACK qS342PSF 6 ~ c = @ ~NL'SEALED 8 ~ ~ 1 0 Direct Thrust Effects Removed a,=@@ NosaGearOff N F ~ ~ 0 2700 RPM 0.3 0.2 8 O-l N Ly 2 0 u -0.1 -0.2 -0.3 -0.4 4 0 4 8 12 16 20 24 Angle of Attack, a - deg OP7O-0022.142 AIcIcR4C'I COklploNV 8-166 MDCA4318 FIGURE 8.6-13 LIFT COEFFICIENT vs ANGLE OF ATTACK, FLOW SURVEY RAKE ON q = 3 4 . 2 PSF 6ic = Oo ~ N L = Oo Inlets Covered 85 = lo Direct Thrust Effects Removed

Run 26 4 = 1 9 6 , = 10°/lOo Nose Gear Off

NF/<= 2700 RPM -4 0 4 a 12 16 2 0 24

Angle of Attack, a - deg

GP76 0 8 2 2 143 McooNNgLL AlRCRAPt COWPANV 8- 16 7 MDCA4318 FIGURE 8.6-14 LIFT COEFFICIENT vs DRAG CCEFFICIENT, FLOW SURVEY RAKE ON q = 34.2 PSF 8 J = lo 6 ~ c = Oo ~ N L = Oo Inlets Covered - . .

Direct Thrust Effects Removed Run 26 = 15O 6 , = 10°/lOo Nose Gear Off N+/BT,f 2700 RPM 1 . 4 1.2 1 .o 0.8 0.4 0 . 2

0 0.02 0.04 0 . 0 6 o .oa 0.10 0.12 0.14 0.1 6

'DAERO OP76 0622 144 8-168 MDC A431 8 FIGURE 8.6-15 PITCHING MOMENT COEFFICIENT vs ANGLE OF ATTACK, FLOW SURVEY RAKE ON q = 34.2 PSF 6 ~ c = Oo ~ N L = Oo Inlets Covered 8J = 10 Direct Thrust Effects Removed .- - ~ - .

Run 26 4 = 15' 6 , = 10°/lOo' Nose Gear On

- . __ .__ . ._ . . - - .

N F ~ = 2700 RPM 0.4 0.3 0.2 0.1 (v

hJ

W a E u -0.1 -0.2 -0.3 -0.4 -4 0 4 8 12 ' 6 20 24

Angle of Attack, a - deg

O P 7 6 . 0 ~ 2 2 746 MDC A431 8 FIGURE 8.6-16 EFFECT OF HORIZONTAL TAIL ON LIFT COEFFICIENT vs ANGLE OF ATTACK

q = 34.2 PSF d ~ c 0' ~ N L = SEALED 8J = 1'

Direct Thrust Effects Removed

6f = 0 ' 6 , - Oo/Oo Nose Gear Off

NF/%= 2700 RPM - - I .- - - - __. - - 1.6 1.4 1.2 1 .o 0.6 0.4 0.2 0 4 8 12 16 20 24 Angle of Attack, a-deg OP78.0622 148 AlCDONNE4.L AdRCRAFT COMPANV 8-176 - . . , MDC A431 8 C O r P 4 a (9 c c MCOONNPLL A#RCRIL\PT COMPANV 8-171 MDC A431 8 FIGURE 8.6-18 EFFECT OF HORIZONTAL TAIL ON PITCHING MOMENT COEFFICIENT vs ANGLE OF ATTACK 4 34.2 PSF 6 ~ c p 0' 6NL = SEALED 8J = 1' Direct Thrust and Ram Air Drag Effects Removed

6f - Oo 6,- Oo/oo Nose Gear Off

N F / ~ = 2700 RPM -4 0 4 8 12 16 20 24

Angle of Attack, cy - deg

aP76.Ob22.148 MDC A4318

8 . 7 AERODYNAMIC LIFT CONFIGURATION - LATERAL-DIRECTIONAL CHAR, - FEZASTICS -. I- -

T h i s s e c t i o n p r e s e n t s t h e 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 and , o n t r o l charac- teristics of t h e aerodynamic l i f t c o n f i g u r a t i o n . S t a b i l i - t y d a t a are presented f o r a n g l e s of a t t a c k o f 0", e", and 16'.

Lateral and d i r e c t i o n a l c o n t r o i e f f e c t i v e - n e s s d a t a are presented f o r a n g l e s of a t t a c k up t o 24'. The s t a b i l i t y d a t a are presented i n both dimensional and c o e f f i c i e n t form. The dimensional form h a s been of t h e p r o p u l s i o n system c o n t r i b u t i o n included i n o r d e r t o emphasize t h e magnitude t o t h e balance-measured f o r c e s and moments. A l l d a t a are presented f o r s t a b i l i t y axes.

Dimensional Data are p r e s e n t e d i n F i g u r e s 8.7-1 The l a t e r a l - d i r e c t i o n a l dimensional d a t a The d i r e c t p r o p u l s i o n system c o n t r i b u t i o n t o t h e measured s i d e through 8.7-9.

f o r c e is t h e r e s u l t o f t h e ram d r a g z o n t r i b u t i o n s of t h e gas g e n e r a t o r s and l i f L / cruise u n i t s . S i n c e t h e ram drag i s e s s e n t i a l l y c o n s t a n t through a n a n g l e of a t t a c k of 16O, t h e p r o p u l s i o n system c o n t r i b u t i o n t o t h e s i d e f o r c e remains c o n s t a n t through t h i s angle. The p r o p u l s i o n system cor.' A b u t i o n s t o t h e measured yawing moment and r o l l i n g moments are t h e r e s u l t o f b o t h t h e d i f f e r e n t i a l t h r u s t bet,Jeen t h e two l i f t / c r u i s e u n i t s and t h e ram drag c o n t r i b u t i o n s from t h e gas g e n e r a t o r s and l i f t l c r u i s e units. The d a t a i n d i c a t e t h a t t h e p r o p u l s i o n system c o n t r i b u t i o n was n e g l i g i b l e except f o r t h e yawing moment a t 16' a n g l e o f a t t a c k .

C o e f f i c i e n t Data The c o e f f i c i e n t form of t h e aerodynamic f o r c e and moment d a t a i s p r e s e n t e d i n F i g u r e s 8.7-10 through 8.7-12. These d a t a are t h e c o e f f i c i e n t form of t h e t o t a l aerodynamic f o r c e and moments p r e s e n t e d i n F i g u r e s 8.7-1 through 8.7-9.

The c h a r a c t e r i s t i c s are e s s e n t i a l l y l i n e a r w i t h s i d e s l i p a n g l e f o L a n g l e s A t 16' a n g l e of a t t a c k , t h e c h a r a c t e r i s t i c s v a r y i n a of a t t a c k of 0 ' and 8'.

n o n l i n e a r manner t y p i c a l c f p o s t - s t a l l o p e r a t i o n . The c o n f i g u r a t i o n does e x h i b i t p o s i t i v e d i r e c t i o n a l s t a b i l i t y , Cn , through 16' a n g l e of a t t a c k . The d i h e d r a l B e f f e c t , C", is s t a b l e €or a n g l e s o f a t t a c k of 8" and 1 5 - , and n e u t r a l l y s t a b l e 0".

f o r L a t e r a l - D i r e c t i o n a l Control E f f e c t i v e n e s s The l a t e r a l - d i r e c t i o n a l c o n t r o l e f f e c t i v e n e s s is presznted i n F i g u r e s 8.7-13 A i l e r o n c o n t r o l power is presented f o r a s i n g l e a i l e r o n d e f l e c t i o n througn 8.7-11.

Dual a i l e r o n d e f l e c t i o n , -25"/+25', d a t a from -25' t o +25O a t 0' a n g l e of attack.

The v a r i a t i o n of rudder e f f e c t i v e - a r e presented as a f u n c t i o n of a n g l e of a t t a c k .

ness w i t h a n g l e o f a t t a c k is present.?d f o r a rudder d e f l e c t i o n of 23'.

MCmONNH4.L. AIRCRAFT COMPANY 8-173 MDC Ab318

- The aileron effectiveness for a single aileron deflec-

Aileron Effectiveness tion from -25" to +2S0 and dual aileron, -25"/+25", deflection is presented in Figure 8.7-13 at 0" angle of attack. The single aileron deflection data exhibit fne trends associated with wings having a supercritical airfoil section in that deflections. The data for the dual the roll effectiveness is higher for the TEU aileron, -25"/+25" as a fv-nction of angle of attack, are presented in Figure 8.7-14.

Aileron effectiveness is maintained at a high level below 12" angle of attack. At higher angles the roll effectiveness is reduced but the ailerons remain effective 28' angle of attack.

through

Rudder Effectiveness - The variation in rudder effectiveness with angle of

attack is presented in Figures 8.7-15 through 8.7-16 for a rudder deflection of 2 3 ' TEL. Rudder effec:iveness is essentially constant below 24" angle of attack.

MCDONN6l.L AIRCRAPS C3MPANV 8-174 M DC A43 1 8 FIGURE 8.7-1 SIDE FORCE vs ANGLE OF SIDESLIP, cy = Oo bH = Oo q = 3 4 . 2 PSF b ~ c = OO ~ N L = SEALED 8 J Io Graphical Summary of Measured and Calculated Force Data Run 129 6f = Oo 6, = Oo/Oo Nose Gear Off NF1-m 2700 RPM a g 4 c a l IL

G

iij -8 -12 -16 -20 -24 -1 2 -8 -4 0 4 8 12 16

Angle of Sideslip. 3 - deg

GP7CO622.192 MCDONNELA AIRCRAFT COMPANY 8 - l i 5 MOCA4318 FIGURE 8.7-2 YAWING MOMENT vs ANGLE OF SIDESLIP. a = 00 6" 10'' q = 341 PSF ~ L C = 0 ' ~ N L p SEALED t ) ~ 0 lo Graphical Summary of Measured and Calculated Moment Data Run129 a,*@ a,=@/@ NoteGearOff N & K ' 2700 RPM Stabilitv Axes Total Aarodynamic Y m m q Moment 4 3 CI c

8 c

E 2 W > ?

-1 -2 -12 -8 -4 0 4 8 12 16

Angle of Sideslip,$ - deg

om6 0622 193 MCDONNELL A-AFT COMPANY 8-176 MDCA4310

B

... .c

Q c

r p c - 4 f P .- e K - 8 -12 -16 -20 -12 -a -4 0 4 8 16

Angle of Sideslip. 3 - deg

OP79.0622~194 MDCA4318 FIGURE 8.74 SIDE FORCE GANGLE OF SIDESLIP, a = 8 '

6~ 0 ' q = 3 4 . 2 PSF b ~ c = 0 ' 6NL = s E k D 8J = 1 '

Graphical Summary of Measured and Calculated Force Data Run 130 6f = Oo 6, * Oo/Oo Nose Gear Off NF/-= 2700 RPM a = 4 c LL

4 - 8

r n -12 -1 6 -20 -24 -1 2 -8 -4 0 4 8 12 16 Angle of Sideslip, - deg OP70.0022 7 95 MDCA4318 FIGURE 8.7-5 YAWING MOMENT v t ANGLE OF SIDESLIP, Q = 8O

6~ * Oo q = 34.2 PSF i5~c = 0 ’ ~ N L = SEALED 8 J = lo

Graphical Summary of Measured and Calculated Moment Data Run 130 5 f 0” sa = Oo/Oo Nose Gear Off N F / G ~ 5 2700 RPM -12 -8 -4 0 4 8 12 16

Angle of Sideslip. J - deg

GP76.0IZZ.t 96 M-NELL AIRCRAFT COMPANY 8-179 MDC A431 8 FIGURE 8.7-6 ROLLING MOMENT vs ANGLE OF SIDESLIP, Q = 8O

6~ = oo q = 34.2 PSF b ~ c = oo ~ N L = SEALED BJ 1'

Graphical Summary of Measured and Calculated Moment Data Pun 130 6f oo p Oo/Oo Nose Gear Off NF/\IBTo 2700 RPM -4

' L Mlasured Rolling Moment

I I

I

I -8 I I I 1 I -12

I I i l

-16

I j j

I I / 1 1 I

-20

i ,i,.i

i

MCOONNELI AIRCRAFT COMPANY 8-180 MDC A431 8 FIGURE 8.7-7 SIDE FORCE vs ANGLE OF SIDESLIP, a = 16' 6~ = 0' q = 34.2 PSF 6 ~ c = 0 ' ~ N L = SEALED 8 J = lo Graphical Summary of Measured and Calculated Force Data

Run 131 6f = 0 ' 6, - Oo/Oo Nose Gear Off

NF/<a 2700 RPM g - 4 c -1 2 -1 6 -20 -24 -12 -8 -4 0 4 8 12 16 Angle of Sideslip, j . deg GJ70 0022 198 MCDONU6U. AIRCRAFT COMPANY

a-iai

MDC A431 8 FIGURE 8.7-8 YAWING MOMENT vs ANGLE OF SIDESLIP, a = 16O 6 1 - 1 = Oo q = 34.2 PSF 6 ~ c = 0' ~ N L = SEALED 8J = 1' Graphical Summary of Measured and Calculated Moment Data Run 131 6 f 0' 6 = Oo/Oo Nose Gear Off NF/- = 2700 RPM -1 2 -8 -4 0 4 8 12 16 Angle of Sideslip, $ - deg 0~7e-oez2 is9 WCD0NNEL.A AIRCRAFT COMPANV 8-182 MDC A431 8 FIGURE 8.7-9 ROLLING MOMENT vs ANGLE OF SIDESLIP, a = 1 6 O b~ = Oo q 34.2 PSF 6 ~ c Oo ~ N L = SEALED 8 J = 1' Graphical Summary of Measured and Calculated Moment Data Run 131 bf = 0 ' 6 , = Oo/Oo Nose Gear Off N F / % ~ 2700 RPM -1 2 -8 -4 0 4 8 12 16

Angle of Sdeslip. 3 - deg

OP76.0622 2 0 0 MCDONNELL A IRCRA P T COMPANY 8-183 MDCA4318 FIGURE 8.7-10 EFFECT OF ANGLE OF ATTACK ON SIDE FORCE COEFFICIENT vs ANGLE OF SIDESLIP 6 ~ ~ 0 ~ qX34.2PSF S L C = O ~ ~ N L = S E A L E D B j = l 0 Direct Thrust Effects Removed 6f = 0 ' 6, = Oo/Oo Nose Gear Off NF/% 1 2700 RPM I I 1 I I

I I I ! i

4 8 12 16 -12 -a -4 Angle of Sideslip. J . deg G P ~ O I Z I 1 0 1 MCDONNELL AIRCRAFT COMPANY 8-184 MDCA4318 FIGURE 8.7-11 EFFECT OF ANGLE OF ATTACK ON YAWING MOMENT COEFFICIENT vs ANGLE OF SIDESLIP SH = 0 ' q 34.2 PSF 6 ~ c = 0 ' ~ N L = SEALED b J = 1' Direct Thrust Effects Removed

5f - 0 ' 6, = Oo/Oo Nose Gear Off

N F / % ' 2700 RPM 0.028 0.024 0.020 0.016 0.012 0.008 0.004 -0.004 -1 2 -8 -4 0 4 8 12 16 Angle of Sideslip. 4. deg OP76 0622 202 WCDONNELLA8HCHAFT COMPANY 8-185 MDC A431 8 FIGURE 8.7-12 EFFECT OF ANGLE OF ATTACK ON ROLLING MOMENT COEFFICIENT vs ANGLE OF SIDESLIP 5~ = 0' q = 34.2 PSF b ~ c = Oo ~ N L = SEALED 8 J = lo Direct Thrust Effects Removed

6f = 0 ' ba - Oo/Oo Nose Gear Off

NFj- 2700 RPM 0.08 0.06 0.04 0.02 w a

0 " -0.02

- 0 . 0 4 -0.06 -0.08 -0.10 - 1 2 -8 -4 0 4 8 12 16

Angle of Sideslip, 4 - deg

OP76 0622 203 MCDONNRLL AIRCRAFT COMPANY 8-186 MDC A431 8 FIGURE 8.7-13 EFFECT OF AILERON DEFLECTION ON LATERAL-DIRECTIONAL CHARACTERISTICS, CY = 3 O ~ H ' O O qp34.2PSF ~ L C = O O ~ N L = S E A L E D O j = l 0 Measured Gata 6f = 0 ' Nose Gear Off N ~ 1 - m 2700 RPM 0.06 0.04 0.02 0 ;j -0.c2 4.04 -0.06 -20 -10 0 10 20 30 (TED) Left Aileron Deflection, S a L . deg fcF7 4 0 6 7 2 200 MCmONNS4.L AIUCRAPT COMPQNY 8-157 MDCA4318 FIGURE 8.7-14 EFFECT OF AILERON DEFLECTTONX~J LATERAL-DIRECTIONAL CHARACTERISTICS, -- 6 -- = 0 ~ __ - - - - -

6~ = 00 q 34.2 PSF 6 ~ c = 0 ' ~ N L = SEALED d J = 1 '

Direct Thrust Effects Removed 6f = 0 ' Sa = -3/:2@ Nose Gear Off w a K u .

a u ' LO 24 23 32 0 4 12 16 Angle of At . , a . deg GP7@.0622 r M M DC A43 1 8 FIGURE 8.7-15 EFFECT OF RUDDER DEFLECTION ON SIDE FORCE SH = Oo q 34.2 PSF 6 LC = 0 ' SNL = SEALED 8 J = lo Direct Thrust Effects Removed 6f = c1' 6 , = Oo/Oo Nose Gear Off NF/% 5 2700 RPM 0 4 8 12 16 20 24 28 32 Angle of Attack, a . Deg w 1 6 - 0 6 2 1 as0 MDC A431 8 FIGURE 8.7-16 EFFECT OF RUODER DEFLECTION ON YAWING AND ROLLING MOMENTS 6 ~ 1 0 ' qS34.2PSF 5 ~ ~ x 0 ' ~ N L = S E A L E D 0 j = l o Direct Thrust Effects Removed

6f - 0 ' 6, = Oo/Oo Nose Gear Off

N F / K p 2700 RPM 0 4 8 12 16 20 24 28 32 Angle of Attack. I . & g OP76 0622 251 MCDONNELI AJRCCIAFT C O R tPANV 8-19G MDC A4318 8 . 8 A I R IXDUCTION SYSTEX PERFORMXVCE The a i r i n d u c t i o n system perf:rmance c h a r a c t e r i s t i c s were measured a t s e l e c t e d test c o n d i t i o n s throughout t h e wind t u n n e l test program on t h e l e f t l i f t i c r u i s e f a n i n l e t , t h e nose l i f t f a n i n l e t , t h e l e f t l i f t / c r u i s e gas g e n e r a t o r i n l e t , and t h e nose f a n engine i n l e t / d u c t . The i n l e t performance rake geonet-:ies u t i l i z e d t o measure t h e i n l e t d a t a were p r e s e n t e d and d i s c u s s e d i n S e c t i o n 3 . The t o t a l p r e s s u r e recovery and s t e a d y state d i s t o r t i o n were measured t o show t h e e f f e c t s of mass flow r a t i o , a n g l e o f a t t a c k , a n g l e o f s i d e s l i p , and forward speed on t h e The r e s u l t s of t h o s e i n l e t p e r f o r s n c e tests are performance of each i n l e t .

d i s c u s s e d for each i n l e t i n t h e s e c t i o n s t h a t follow.

L e f t L i f t / C r u i s e Fan I n l e t The e f f e c t s of a n g l e of a t t a c k and f a n speed on t h e performance of t h e over- the-wing mounted l i f t i c r u i s e f a n i n l e t are p r e s e n t e d i n F i g u r e 8.8-1. The average t o t a l p r e s s u r e recovery ( P T ~ / P T ~ ) and t h e i n l e t d i s t o r t i o n f a c t o r (PThigh - PTlou/ P T ~ ~ ~ ) are p r e s e n t e d along w i t h t h e i n l e t mass flow r a t i o (..%/Am) as a f u n c t i o n of c o r r e c t e d f a n speed a t t h e h i g h e s t t u n n e l speed ( V , = 103 k n o t s ) t e s t e d . The mass flow r a t i o d a t a shown w a s c a l c u l a t e d u s i n g t h e f a n e x i t r a k e , as p r e v i o u s l y Jes- c r i b e d . The d a t a p r e s e n t e d i n t h e f i g u r e uere measured w i t h t h e model i n t h e c r u i s e mode c o n f i g u r a t i o n , nose ur?it o f f .

A t a mass flow r a t i o of 1.0, t h e i n l e t p e r f o r a a n c e i s snown t o m a i n t a i n a very Above t h i s up K O ar-gles of 32" it con- high l e v e l up t o an a n g l e of a t t a c k of 20".

t i n u a l l y d e i e r i o r a t e s . This f a l l o f f i n perfornance a t t h e h i g h e r 3 is due t o t h e u n s t e a d i n e s s and t h e subsequent s e p a r a t i c n o f t h e f l c w on t h e i n b o a r i p a n e l as 3 is i n c r e a s e d . Ncte t h a t a t t h e lower a s s flcw r a t i o s , i.e., reduced 7ower levels, t h e perf?mnance d e t e r i o r a t e s a t lower a n g l e s of a t t a c k . This is alsd due t o Iriboard w i n , panel s e p a r a t i o n , wSich is o c c u r r i n g a t lower 2 due t o a lesser amount o f i n l e t induced flow over t h e inboard wing. Thrcu5hout t h e t e s t , t h e t u f t s on t h e wing were observed t o i n d i c a t e flow s e p a r a t i o n a t a n g l e s o f a t t a c k of a p p r o x i n a t e l y 8' h i g h e r on t h e inboard wing ahead o f t h e i a l e t than on :>.e m t b o a r d wing, a t mass flow r a t i o s g r e a t e r than 1.0. TIie l i f t / c r n i s e i n l e t ra!:e d a t a shown h e r e h s u b s t a n t i a t e t h e s e o b s e r v a t i o n s .

It should be noted h e r e t h a t tnroughout t h e t e s t p r o a r a n a t iower a n g l e s o f a t t a c k (Oo-200) and h i g n e r mass flow r a t i o s (.&,/Am ). 1.0) t h e i n l e t r p - ~ . * r y ransed Setween 0.999 and 1.00. The r e a s c n f o r t h e s e very nigh levels o f peerLor?lance i s t h e low s p e c i f i c flow (:;F.L~/A) of t h e low far, p r e s s u r e r a t i o X3768 t u r b o t i p far.

(0.35 and below). The i n t e n t of and t h e r e s u l t i n g low i n l e t t h r o a t Xach nuribers MDC A4318 the i n l e t performance d a t a p r e s e n t e d h e r e i n f o r a l l i n l e t systems i s t o show treris and t h e s p e c i f i c e f f e c t s of a n g l e o f a t t a c k , s i d e s l i p , and mass flow r l t i o , and n o t n e c e s s a r i l y t h e a b s o l u t e levels of perfornance, f o r example, o f h i g h e r p r e s s u r e ratic f a n systems.

- % e e f f e c t o f s i d e s l i p angle on i n l e t p e r f o r n a n c e a t v a r i o u s a n g l e s of e t t a c k is shown i n F i g u r e s 8.8-2 and 8.5-3 f o r t h e range of test c o n d i t i o n s where i n l e t d a t a were measured.

S i d e s l i p a n g l e s of up t o +12O ( l e f t i n l e t on leeward s i d e of f m e l a g e ) shcw no e f f e c t on l i f t / c r u i s e i n l e t performance a t a n g l e s of a t t a c k of 12" and lower.

The d e t e r i o r a t i o n i n performance a t h i g h e r a z g l e s of a t t a c k and s i d e s l i p is a p p a r e n t in t h e f i g u r e s .

I n l e t d i s t o r t i o n levels were found t o b e less than 10% throughout a l l test c o n d i t i o n s d u r i n g t h i s program.

The e f f e c t s o t forward ;peed on i n l e t performance v a r i a t i o n are shonn i n Figure 8.8-4.

Only t h e d i s t o r t i o n w a s found t o show any v a r i a t i o n , and t h i s occurred only at t h e h i g h e r a n g l e s of a t t a c : .

A summary of t h e l i f t / c r u i s e f a n i n l e t p e r f o r m a t c e i s p r e s e n t e d i n F i g u r e 8.8-5.

The d a t a are p r e s e n t e d i n terms of t h e nondinensional i n l e t loss c o e f f i c i e n t High speed (?io = 0 . 5 t o 0.9) p s r f o m n c e d a t a (lPT/qTH) for c o r r e l a t i o n purposes.

f o r t h i s i n s l t are contained i n Reference (3).

L e f t L i f t / C r u i s e Gas Generator I n l e t The e f f e c t s of a n g l e o f a t t a c k and e n g i n e speed on t h e performance of t h e f u s e l a g e s i d e mounted gas g e n e r a t o r i n l e t a r e snwn i n F i g u r e 8.8-6. The i n l e t are p r e s e n t e d i n t h e same manner as t h e far. i n l e t d a t a d e s c r i b e d above. The d a t a e f f e c t s of both a n g l e of a t t a c k and s i d e s l i p a r e shown i n F i g u r e s 8.8-3 and 8.8-7.

In g e n e r a l , t h e i n l e t e x h i b i t e d very good perfornance tfiroughout t h e test range of T o t a l p r e s s u r e recovery w a s found t o d e t e r i o r a t e less t h a n 1 8 5 , and v a r i a b l e s .

d i s t y r t i o n w a s measured t o be less than 5 2 throughout t h e a n g l e of a t t a c k and s i C e s l i p a n g l e s t e s t e d .

Nose L i f t Fan I n l e t The e f i e c t s o f a n g l e of a t t a c k and fan speed on t h e performance of t h e n o s e f a n i n l e t f o r t h r e e t u n n e l v e l o c i t i e s are ? r e s e n t e d i n Ffgures 8.3-8, 8.8-9, and The d a t a are shown f o r 3's ranging from -4" t o +20° which covered t h e 3.8-10.

l i f t test range of t h i s program. As shown i n t h e f i g u r e s , t h e e f f e c t s o f powered a n g l e of a t t a c k up t o 20" f o r all forward speeds had very l i t t l e e f f e c r cn e i t h e r The d a t a psesented a l s o t h e t c t a l p r e s s u r e recovery o r d i s t o r t i o n o f t h e i n l e t , cover t h e f a n s p e 4 r a c g e s t e s t e d i n t h e powered l i f t mode.

The o p e r a t i n g mass i n l e t is p r e s e n t e d i n term of t h c n a s s flow r a t i o (.&/i;a~> r a t i e r flow of t h e than t h e more t y p i c a l i n l e t v e l o c i t y r a t i o ( V o , " . ' ~ ~ ) f o r reasons of conparison and MCDONNELL AIRCRAFT COMPANY 8-192 YDC A4318 c o n s i s t e n c y w i t h t h e l i f t / c r u i s e i n l e t performance d a t a p r e v i o u s l y d i s c u s s e d . x p l o c of t h e i n l e t v e l o c i t y r a t i o ! Y 0 / V ~ ~ ) expressed i n terms of t h e mass flow r a t i o (A.,/AHL> f o r t h e nase f a n i n l e t i s shown i n F i g u r e 3.8-11.

The e f f e c t s of s i d e s l i p a l s o had very l i t t l e i f no e f f e c t s on i n l e t recovery and d i s t o r t i o n , as shown i n F i p r e 8.8-13. The d a t a presented a l s o encompass t h e test v a r i a b l e s exaninert i n t h e powered l i f t mode.

range of X sunma,y of t h e nose f a n i n l e t performance is p r e s e n t e d i n Figure 8.8-13.

The recovery d a t a h e r e are p r e s e n t e d i n term of t h e nondimensional i n l e t l o s s c o e f C i c i e n t ('PT/qTH) v e r s u s t h e i n l e t v e l o c i t y r a t i o f o r c o r r e l a t i o n purposes.

' :3::a p r e s e n t e d i n t h i s f i g u r e encompass t h e complete range of t u n n e l and f a n - - i d s r e s t e d on t h i s i n l e t . For t h e nominal o p e r s t i n g range o f t h e nose fan z l e t f o r a V/STOL a i r c r a f t of t h i s type, t h e performance of t h i s i n l e t e x h i b i t e d as shown i n t h e f i g u r e .

e x e l l e n t c h a r a c t e r i s t i c s , NK=DCWNELL AIRCRAFT COMPANY 8-193 MDC A431 8 FIGURE 8.8-1 EFFECTS OF ANGLE OF ATTACK AND FAN SPEED ON LIFT/CRIJISE FAN INLET PERFORMANCE U H Inlet V, = 103 K h I I c 6 - - .

, I , ‘ L L ~ !

! I I I

I - --. -- - . -

1 b 12 16 20 24 28 32 Angle of Attack, G oeg GP70 0 6 2 1 3 1 MCmONNELL AlRCRAPT COMPANV 3-194 I% DC A43 1 8 FIGURE 8.8-2 EFFECTS OF ANGLE OF ATTACK AND SIDESLIP ON LIFT/CRUISE FAN INLET PERFORMANCE L/H Inlet - -_I--.

V, = 103 Kts NF/* = 2700 RPM

- - P n

Y k u Y I

I 1 % . 1 I

Angle of Atrack, Q . deg MCOONNaLL AIRCRAFT COMPLINY 8-195 MDCA4318 FIGURE 8.8-3 EFFECTS OF SIDESLIP ANGLE AND ANGLE OF ATTACK ON LIFT/CRUISE FAN AND GAS GENERATOR INLET PERFORMANCE UH Inlets V, 103 Kts N F I ~ ; = 2700 RPM t a \ (v c a .

\ (v Inlet Distortion , - 4 0 4 8 12 16 20 24 GP7O 0622 38 Angle of Si+sllp, J . deg MCDONNEkL A8RCHAPT COMPANY 8-196 MDC A431 8 FIGURE 8.84 EFFECTS OF FORWARD SPEED AND ANGLE OF ATTACK ON LIFT/CRUISE FAN INLET PERFORMANCE L/H Inlet NF/v% = 3600 RPM MCDONNRLL AtRCRAP 7 COMPANY 8-197 MDC A43 1 8 FIGURE 8.8-5

I

LIFT/CRUISE FAN INLET PERFORMANCE SUMMARY NondimensionalizedTotal Pressure Loss Data V, = 103 Knots 0 0.4 0.8 1.2 1.6 2.0 2.4 2.8

Inlet Velocity Rat:o - V,,’VTH

GP70 0 6 ; 2 281 MCOONNPLL AIRCRAFT COMPANY 8- 19 8 MDC A431 8 FIGURE 8.8-6 EFFECTS OF ANGLE OF ATTACK AND ENGINE SPEED ON LIFT/CRUISE GAS GENERATOR INLET PERFORMANCE L/H fnlet V, = 103 Kts

1 +-

; I , ; I I I

: I

20 24 28 32 0 4 8 12 16 Angle of Attack, a . deg G P 7 @ 0 6 2 2 11 MCk30NNELL. AIRCRACY C3MPANV 8-1 99 MDC A431 8 FIGURE 8.8-7 EFFECTS OF ANGLE OF ATTACK AND SIDESLIP ON LIFT/CRUISE GAS GENERATOR INLET PERFORMANCE L/H Inlet

V, - 103 Kts

N ~ l f i ~ 5 2700 RPM MCDONNHLL AIRCHAP T COMPANY 8-20U MDC 4431 8 FIGURE 8.8-8 EFFECTS OF ANGLE OE ATTACK AND FAN SPEED ON NOSE FAN INLET PERFORMANCE V, = 47 Kts I Note. Fan weoar shown cover range restad I -4 U 4 8 12 16 20 24 Angle of Attack, . deg GP76 (1621 39 MCD0NNSl.L AlRCRAPT COMPANY 8--31 MDCA4318 FIGURE 8.8-9 EFFECTS OF ANGLE OF ATTACK AND FAN SPEED ON NOSE FAN INLET PERFORMANCE V0=63Kts

Angle of Attack, a - deg

GP70 0822-41 MCDONNELL AIRCRIICI COMPANV 8-202 PADC A431 8 FIGURE 8.8-1C EFFECTS OF ANGLE OF ATTACK AND FAN SPEED ON NOSE FAN INLET PERFORMANCE V,=78 K n -4 4 8 12 16 20 24

Angle of Attack, Q - deg

GP7B-0621.40 IWC0ONNEl.L AIRCRAFT COMPANY 8-203 MDC A431 8 FIGURE 8.8-11 NOSE LIFT UNIT INLET VELOCITY RATIOS 1.4 " 0 0 . 2 0.4 -0.6 0 . 8 1 .o 1.2 I Inlet Velocity R&& V,/VTH GP76.0622~294 MCDONNELL A~RCRAPT COMPAP r 8-204 MDCA4318 FIGURE 8.8-12 EFFECTS OF SIDESLIP ANGLE ON NOSE FAN INLET PERFORMANCE a=OO VO=103Kts I

Sideslip Angle, f l - deg

OP76-0122-293 MCDONNPLL AIRCRAFT caMPaniv 8-205 MDCA4318 FIGURE 8.8-13 NOSE FAN INLET PERFORMAYCE SUMMARY Nondimensionalited Total PressJre Loss Data a m 0 0 0 0.4 0.8 1.2 1 T 6 2 .o 2.4 2.8

Inlet Velocity Ratio - V,/VTH

OP76-062242 m N E U AIRCRAFT COWPANV 8-206 MDC A4318 9. OUTSIDE STATIC TEST RESULTS , The r e s u l t s of t h e o u t s i a e s t a t i c test program conducted on t h e l a r g e scale powered model are p r e s e n t e d i n t h i s s e c t i o n . me d i s c u s s i o n s cover t h e p r o p u l s i o n system c a l i b r a t i o n s , t h e ground e f f e c t s on t o t a l i n s t a l l e d l i f t , t h e ground e f f e c t s on i n l e t r e i n g e s t i o n , and t h e flow v i s u a l i z a t i o n tests.

9.1 PROPULSION SYSTEM CALIBRATIONS I n d i v i d u a l Unit C a l i b r a t i o n s The balance-measured s t a t i c gross t h r u s t and t h e rake-measured i d e a l gross t h r u s t s are p r e s e n t e d as a f u n c t i o n of f a n speed squared f o r t h e l e f t and r i g h t The r e s u l t s are shown f o r l i f t l c r u i s e u n i t s i n Figures 9-1 and 9-2, r e s p e c t i v e l y .

t h e 9 0 ' v e c t o r p o s i t i o n only, as t h i s was t h e o n l y v e c t o r p o s i t i o n t e s t e d on t h e 4 0 ' x 80' wind tunnel lift/cruise units. For comparison purposes, r e s u l t s of t h e s t a t i c c a l i b r a t i o n s f o r each u n i t are a l s o shown i n t h e f i g u r e s . F a i r l y good exists between t h e two t e s t programs, p a r t i c u l a r l y on t h e r i g h t u n i t . A agreement rear v i e w of t h e model showing t h e l i f t / c r u i s e v e c t o r i n g units is given i n Fig- ure 9-3. F i g u r e 9-4 p r e s e n t s t h e balance- and rake-measured t h r u s t s f o r t h e f o r - ward l i f t u n i t a t a geometric v e c t o r angle of 95'.

Comparisons between t h e t u n n e l measured and s t a t i c test measured n o z z l e t h r u s t f o r t h e l e f t , r i g h t and nose l i f t u n i t s are p r e s e n t e d c a l i b r a t i o n c o e f f i c i e n t s (CF) As shown i n t h e f i g res, good agreement e x i s t s between t h e i n Figures 9-5 and 9-6.

two l i f t / c r u i s e u n i t c a l i b r a t i o n s ; however, a s h i f t of approximately 5% w a s measured Comparisons of t h e r e s u l t a n t t h r u s t v e c t o r angles f o r each u n i t on t h e nose u n i t .

are p r e s e n t e d i n F i g u r e 9-7 and 9-8, showing good agreement w i t h t h e 40' x 80' tunnel s t a t i c c a l i b r a t i o n s . The a b s o l u t e l e v e l of accuracy of t h e o u t s i d e s t a t i c b a l a n c e and rake c a l i b r a t i o n d a t a is of lesser importance t o t h e v a l i d i t y of t h e s e test r e s u l t s than t h e v a r i a t i o n i n l e v e l measured f o r each h e i g h t and model con- f i g u r a t i o n t e s t e d . Thzse comparisons w i t h t h e 40' x 80' d a t a are p r e s e n t e d h e r e i n f o r r e f e r e n c e purposes and as a check on t h e o u t s i d e test s e t u p .

Fan Performance Map The X376B t u r b o t i p f a n performance map is p r e s e n t e d i n c o r r e c t e d form i n Figures 9-9 and 9-10. The n o z z l e t o t a l p r e s s u r e r a t i o i n F i g u r e 9-9 and the i d e a l gross t h r u s t i n FFgure 9-10 are shown v e r s u s t h e t o t a l f a n - p l u s - t i p - t u r b i n e a i r f l o w as a f u n c t i o n of v a r i o u s fan speeds f o r t h e t h r e e a l t e r n a t e n o z z l e exhaust areas t e s t e d . The values i n t h e performance map were c a l c u l a t e d using the f a n and t i p .

MCOONNRLL AIRCRAFT COMPANY 9-1 MDC A4318 turbine exit rake data and are presented directly as measured. The fan performance maps are used in evaluating the individual behavior of the three lift units during operation in ground effects. The effects of both nozzle back pressure and inlet hot gas reingestion on fan thrust and airflow rate can be assessed utilizing these fan maps.

MCDONNELL AIRCRAFT COMPANY 9-2 MDCA4318 FIGURE 9-1 LEFT LIFT/CRUISE UNIT CALIBRATION RESULTS Static Test Calibrations 6 LC = goo - -

-. -- . __ - - -

- Corrected Fan Speed Squared, [ N ~ I f l y , l - (rpmI2 x op1o.oe22.2~2 M C - U AIRCRAFT COMPIINY 9-3 MDCA4318 FIGURE 9-2 RIGHT LIFT/CRUISE UNIT CALIBRATION RESULTS Static Test Calibrations

s LC = 900

Corrected Fan Speed Squared, [NF/fio] - (rprn)’ x

OP76.0622.291 MCDOUNELL A8RCRAPT COMPII\NV 9-4 MDC A431 8 FIGURE 9-4 NOSE LIFT UNIT CALIBRATION RESULTS Static Test Ca I i brat ions b ~ ~ = 9 5 ' 1 200 o_ Rake Measured ideal Thrust ( F I I ~ A ) a

c 1000

u .

U- ul x

+

z

ij 800

s

Balance Measured Static Thrust ( F , i b ~ ) 8 10 - 12 14 6 4 2

- - -

Corrected F a n Speed Squ;--l, [ N ~ ~ f i ~ l . (rprnI2 x o~m-oezz 290 MCOONNaLL AIRCRAFT COMPANV 9-6 MDCA4318 FIGURE 9 5 LIFT/CRUISE UNIT THRUST COEFFICIENT bOMPARISON Propulsion System Calibration Results v, P O 20 40 60 80 1 0

Geometric Deflection Angle, 6 LC - deg

OP76-0622.306 MCD0NNEL.L AIRCRA P T COMPANY - 7 MDC A43 1 8 FIGURE 9-6 NOSE LIFT UNIT THRUST COEFFICIENT CUMPARISON Prop3lsion System Cali brati on R esu I t s W , 5 0 Geometric Deflection Angle, ~ N L . de3 OP70 0622.501 MCOONNELI. AIRCRAFT COMPANY 9-8 MDC A43 1 8 FIGURE 9-7 LIFTKRUISE UNIT THRUST VECTOR ANGLE COMPARISON Propulsion System Calibration Results

Geometric Deflection Angle, 6 LC - deg

GP76-0122.303 MDCA4318 FIGURE 9-8 NOSE LIFT UNIT THRUST VECTOR ANGLE COMPARISON Propulsion System Calibration Results v,=o !O

Geometric Deflection Angle. ~ N L - deg

QP760622.302 MCDONNRLL AIRCRAFT COMPL\NY 9-10 MDC A43 1 8 MDCA4318 MCOONNEU AIRCRAFT COMPANY 9-12 MDC A4318 9 . 2 GROUND EFFECTS ON AIRCRAFT LIFT LOSS c , The e f f e c t s of alternate ground h e i g h t s and s e l e c t e d model v a r i a b l e s on t h e t o t a l i n s t a l l e d l i f t d u r i n g t h r e e u n i t o p e r a t i o n , along w i t h t h e i n d i v i d u a l u n i t ground e f 5 -ts test results are given i n t h i s s e c t i o n .

A l l d a t a p r e s e n t e d are based m t -easy state measurements o b t a i n e d w i t h t h e Vidar d a t a a c q u i s i t i o n system.

Three U r --- it Operation

The -aslance-measured t o t a l l i f t a t t h e t h r e e ground h e i g h t s t e s t e d i n t h i s program are p r e s e n t e d as a f u n c t i o n of f a n speed i n F i g u r e 9-11. These l i f t d a t a are p l o t t e r v e r s u s t h e average ambient c o r r e c t e d f a n speeds ( r a t h e r than inlet correctec! :an speeds) i n o r d e r t o compare a t each a l t i t u d e t h e combined ground e f f a c t s i x h d i n g i n l e t r e i n g e s t i o n . The d a t a as p r e s e n t e d i n t h e f i g u r e i n c l u a e t h e ecfects of suckdown p l u s f o u n t a i n , f a n back p r e s s u r e , and i n l e t r e i n g e s t i o n as a f f e c t e d by a i r c r a f t ground h e i g h t . As shown by t h e results in t h i s f i g u r e , e s s e n t i a l l y no n e t l i f t l o s s i n ground e f f e c t s w a s measured w i t h t h i s a i r c r a f t test model., p a r t i c u l a r l y a t t h e h i g h e r f a n speeds.

T h c rake-measured i d e a l t h r u s t s are p r e s e n t e d f o r each l i f t u n i t f o r a l t i t u d e s of 21.0, 8.3, and 3.3 f e e t i n Figures 9-12, 9-13 and 9-14, r e s p e c t i v e l y , f o r t h e same test p o i n t s as F i g u r e 9-11. The rake-measured i d e a l t h r u s t s are used t o e v a l u a t e t h e t h r u s t v a r i a t i o n of t h e i n d i v i d u a l u n i t s s e p a r a t e l y from t h e t o t a l n e t i n s t a l l e d l i f t . ..ne i d e a l t h r u s t v a r i a t i o n s shown i n c l u d e t h e e f f e c t s of inlet reing:?stion and back p r e s s u r e , t h e r e b y a l l o w i n g a comparison w i t h t h e f o r c e d a t a t o y i e i d t h e n e t p r o p u l s i o n induced ground e f f e c t s o c c u r r i n g on t h e a i r c r a f t model.

The r e s u l t s of t h e r a k e determined i n d i v i d u a l t h r u s t measurements w i t h a l l t h r e e u n i t s o p z r a t i n g are p r e s e n t e d i n F i g u r e 9-15.

The r e l a t i v e change i n t h r u s t ("/aF versiis t,ie model h e i g h t r a t i o (H/D) are p r e s e n t e d f o r each u n i t a t a s e l e c t e d c o . ected f a n speed of 3600 rpm. The model h e i g h t r a t i o (H/D) f o r t h i s and a l l - irves p r e s e n t e d h e r e i n i s based on t h e average e x i t flow a r e a ( 7 ) of a l l t h r e e l i f t u n i t s (D = 39.1"). The reference t h r u s t used diameter AVG i n t h e * gure and throughout t h e r e p o r t was t h e v a l u e measured a t t h e 21.0 f o o t t h r u s t a t t h e reduced model h e i g h t s mode.. height. The expected d e t e r i o r a t i o n of i s apparent ir. the f i g u r e as t h e e f f e c t s of both i n l e t temperature r e i n g e s t i o n and back preL.:u;s are encountered. rise i n t h e t h r u s t of t h e nose l i f t The sharp u n i t as t t 2 lower ground h e i g h t s a r e reached was found t o be t h e n e t r e s u l t of ' b3th reduced i n l e t r e i n g e s t i o n and f a v o r a b l e f a n bock p r e s s u r e at t h e lowest g x n d h e i g h t . Inspection of t h e nose fan and gas generator i n l e t temperature MCDONNSLL AIRCRAFT COMPANY 9-13 .MDC A431 8 FIGURE 9-11 EFFECT OF GROUND HE'3HT ON TOTAL MEASURED LIFT All Units Operating a = oo 1 200 0 . _ _ 1200 1600 2000 2400 2800 3200 3600 4000 Corrected Fan Speed, N F l q - rprn GP76.0622 1 2 1 M C B N R L L AIRCRAFT COMPANV 9- 14 MDCA4318 FIGURE 9-12 INDIVIDUAL UNIT IDEAL THRUST MEASUREMENTS All Units Operating Model Height = 21.0 Ft 1 200 1200 1600 2000 2400 2800 3200 3600 4000 Corrected Fan Speed, NF/% - rpm GP70-0822 1 2 2 MCDONNBlL AIRCRAFT COMPANY 9-15 MDC A431 8 .

FIGURE 9-13 INDIVIDUAL _-___-____ UNIT IDEAL THRUST MEASUREMENTS All Units Operating - .- Model Height = 8.3 Ft 1 600 1 200 1200 1600 2000 2400 2800 3200 360C 4000 Corrected Fan Speed, N F / V F - rpm OP76 0622 123 MCDONNKLL AIRCRAFT COMPANY 9-16

MDC ~ 4 3 1 a

FIGURE 9-14 INDIVIDUAL UKlT IDEAL THRUST MEASUREMENTS All Units Operating Model Height = 3.3 Ft " 1200 1600 2000 2400 2800 3200 3600 4000 Corrected Fan Speed, N F / G o ~ rpm GP76.0622.121 MCDONNELL AIRCRAFT COMPANY 9-17 MDC A431 8 FIGURE 9-15 EFFECT OF GROUND HEIGHT ON THE INDIVIDUAL UNIT THRUST All Units Operating Correctad Fan Speed 1 3600 RPM 0176.0622-1 30 Model Height Ratio - H/D I i MCOONNaLL AIRCRAFT COMPANY 9- 18 MDC A4318 r e i n g e s t i o n d a t a i n S e c t i o n 9.3 of t h i s r e p o r t e x p l a i n s t h e more f a v o r a b l e r e i n - g e s t i o n e f f e c t s . The i n c r e a s e d back p r e s s u r e a t t h e lower ground h e i g h t caused a r e d u c t i o n i n e f f e c t i v e n o z z l e a r e a .

The nose f a n geometric nozzle a r e a was o v e r s i z e t o s t a r t with, t h e r e f o r e t h e n o z z l e area match was a c t u a l l y improved a t t h e lower ground h e i g h t . T h i s moved t h e f a n o p e r a t i n g p o i n t t o a more f a v o r a b l e l o c a t i o n on t h e f a n map of Figure 9-10 and t h u s increased t h e t h r u s t output. The r e l a t i v e v a r i a t i o n of t h e combined t o t a l t h r u s t w i t h h e i g h t changes is shown by t h e dashed l i n e i n Figure 9-15.

The comp-rison of t h e combined t o t a l rake-measured t h r u s t v a r i a t i o n w i t h t h e t o t a l balance-measured l i f t v a r i a t i o n a t a f i x e d ambient c o r r e c t e d f a n speed of 3600 rpm i s presented i n F i g u r e 9-16.

The n e t balance-measured l i f t change w i t h ground h e i g h t ( l i f t l o s s ) was found t o d e c r e a s e less than 1% over t h e t h r e e h e i g h t s t e s t e d . The t h r u s t v a r i a t i o n ( t h r u s t l o s s ) , however, was found t o d e c r e a s e as much a s 7% a t t h e i n t e r m e d i a t e h e i g h t t e s t e d . It is a p p a r e n t from t h e s e d a t a t h a t a n e t p o s i t i v e induced f o r c e is o c c u r r i n g on t h i s mociel probably due t o f a v o r a b l e f o u n t a i n e f f e c t s . A s was s t a t e d p r e v i o u s l y , both t h e rake-measured t h r u s t and balance- measured t o t a l l i f t have t h e e f f e c t s of r e i n g e s t i o n and Dack p r e s s u r e i n t h e v a l u e s presented, t h e r e f o r e i n d i c a t i n g thaL t h e d i f f e r e n c e s between t h e two measurements w i t h ground h e i g h t are due t o induced r a t h e r than d i r e c t t h r u s t f o r c e s .

The e f f e c t s of gas g e n e r a t o r i n l e t s h i e l d s on t h e r e d u c t i o n of t h e gas g e n e r a t o r i n l e t r e i n g e s t i o n were found t o b e f a v o r a b l e and are discussed in Section 9.3. However, as shown in t h e F i g u r e s of 9-17 through 9-19, t h e e f f e c t s of i n l e t s h i e l d i n g on t h e balance-measured t o t a l l i f t were found t o be v e r y small.

Figure 9-17 shows e s s e n t i a l l y no e f f e c t f o r t h e l a r g e s h i e l d a t a model h e i g h t of 21.0 f t . A small but apparent p o s i t i v e e f f e c t due t o t h e a d d i t i o n of t h e s m a l l s h i e l d a t t h e i n t e r m e d i a t e 8.3 f t a l t i t u d e i s shown i n Figure 9-18. A comparison of t h e l a r g e and small i n l e t s h i e l d s and t h e j r e f f e c t on t o t a l i n s t a l l e d l i f t l o s s is shown i n Figure 9-19 a t t h e lowest h e i g h t of 3.3 f e e t . A s shown in t h e f i g u r e , e s s e n t i a l l y no d i f f e r e n c e e x i s t e d bctweec t h e l a r g e and small s h i e l d s , although a q m a l l b c t n e g a t i v e e f f e c t on l i f t was measured by both s h i e l d s a s compared w,th t h e s h i e l d o f f c o n f i g u r a t i o n . This n e g a t i v e e f f e c t i n d i c a t e s t h a t , although t h e i n l e t r e i n g e s t i o n is reduced and t h r u s t increased ( s e e Section 9 . 3 ) , t h e r e is a s l i g h t l y g r e a t e r l i f t loss d u e t o an apparent i n c r e a s e i n suckdown. Figure 9-20 p r e s e n t s t h e e f f e c t of s h i e l d d e f l e c t i o n a n g l e (6,) on t h e balance-measured t o t a l l i f t a t a c o r r e c t e d f a n s p e e d of 2500 rpm a t t h e 3.3 f o o t model h e i g h t . This d a t a i n d i c a t e s t h a t an optimum dcflect.lon a n g l e e x i s t s a t approximately 45", WCDONNELL A # W C W A F 1' COMPA N V 9-19 MDCA4318 FIGURE 9-16 EFFECT OF GROUND HEIGHT ON TOTAL LIFT AND THRUST All Units Operating Corrected Fan Speed ( N t A / F o ) = 3600 RPM b - 4 - 8 -12 0 1 2 3 4 5 6 7 Modd Height Ratio, H/D OP7C0622~129 MCDONNELL A#RCRLL F T COMPA N V 9-20 V O C A 4 3 1 8 FIGURE 9-17

* '

E F F L T OF INLET SHIELDING ON TOTAL MEASURED LIFT All 'Jnits Operating Model Height = 21.0 F t 0 ) 0 -_ m m MCOONNFLL AIKCRAFT C0MPAV.W 9 - 2 1 MDCA4318 FIGURE 9-18 EFFECT OF INLET SHIELDING ON TOTAL MEASURFD LIFT All Units Operating Model Height = 8.3 Ft P MCDONNgLL AIRCUAPT COMPANY 9 - 2 2 MDCA4318 FIGURE 9-19 EFFECT OF INLET SHIELDING ON TOTAL MEASURED LIFT All Units Operating Modal Height = 3.3 Ft MDC A431 8 FIGURE 9 - Z EFFECT OF SHIELD DEFLECTION ANGLE ON TOTAL MEASURED LIFT All Units Operating Model Height = 3 . 3 Ft N$fi0 = 2500 RPM 20 40 60 80 100 120 Shield Der'lection Angle, 6s deg GP7e.0022-128 lwcooNNaLL A-APT COW-NY 9-24 ?fDC A4318 a p p a r e n t l y providing a c o m p r o d s e between t h e opposing inlet r e i n g e s t i o n and suck-

,'

down e f f e c t s a t t h i s i n t e r m e d i a t e angle.

I n d i v i d u a l Unit Operation The balancc-measured l i f t of t h e i n d i v i d u a l u n i t s a t t h r e e v a l u e s of ground h r i g h t are presentee h F i g u r e s 9-21, 9-22, and 9-23 f o r t h e l e f t , r i g h t , and nose units, r e s p e c t i v e l y . The data are p l o t t e d as a f u n c t i o n of ambient c o r r e c t e d far: I speed, $/'€lT0 . The l i f t loss (or gain) i n ground e f f e c t is m c h more a p p a r e n t f o r t h e s e i n d i v i d u a l u n i t runs than f o r t h e p r e v i o u s l y d i s c u s s e d three unit cow bined runs. The lift/cruise u n i t s showed an i n c r e a s i n g l i f t loss as ground h e i g h t was reduced. The nose l i f t unit showed t h e same l o s s t r e n d as ground h e i g h t s was reduced u n t i l t h e 8.3 f t level &s reached below which t h e l i f t began i n c r e a s i n g .

Inlet r e i n g e s t i o n on t h e i n d i v i d u a l test runs w a s m i n i m a l , as no f o u n t a i n or l o c a l upwash flow was generated. The l i f t loss (or g a i n ) as measured by t h e b a l a n c e d a t a was t h e r e f o r e due t o e i t h e r i n c r e a s i n g f a n back p r e s s u r e , suckdown, o r both, The rake-measured i d e a l gross t h r u s t f o r t h e l e f t , r i g h t , and nose units are p r e s e n t e d i n F i g u r e s 9-24, 9-25, a r d 9-26, r e s p e c t i v e l y , f o r each ground h e i g h t t e s t e d . These d a t a a l s o are p r e s e n t e d versus c o r r e c t e d f a n speed and are from t h e same run as t h e t h r e e balance-measured data f i g u r e s . A l l t h r e e units e x h i b i t e d a r e l a t i v e l y small change in t h e rake-measured t h r u s t when a l t i t u d e w a s decreased, as shown i n t h e f i g u r e s . The left and r i g h t u n i t t h r u s t decayed as ground h e i g h t decreased. However, t h e nose u n i t t h r u s t i n c r e a s e d w i t h reduced ground h e i g h t s .

Once a g a i n , inlet r e i n g e s t i o n played an i n s i g n i f i c a n t r o l e i n t h e va-riation of those -ake-measured t h r u s t d a t a , w i t h t h e t h r u s t v a r i a t i o n e f f e c t s , whether p o s i t i v e or n e g a t i v e , b e i n g a r t r i b u t a d h e r e t o e f f e c t i v e n o z z l e area changing, i.e., back p r e s s u r e e f f e c t s .

Comparisons of t h e balance-measured l i f t loss d a t a w i t h t h e rake-measured t h r u s t loss d a t a are p r e s e n t e d i n Figures 9-27, 9-28, and 9-29 f o r t h e l e f t , r i g h t , and nose l i f t u n i t s , r e s p e c t i v e l y . The d a t a ark p r e s e n t e d as a f m c t i o n of model h e i g h t r a t i o (H/D) f o r a f i x e d c o r r e c t e d €an speed of 3600 pm. The 21.0 f o o t model heiCht is used as t h e r e f e r e n c e a l t i t u d e from which t h e p e r c e n t a g e changes S o t e t h a t b o t h l i f t l c r u i s e u n i t s show a t h r u s t decay maximizing a t are computed.

The n e g a t i v e d i f f e r e n c e between between 3 and 4 X , a t t r i b u t a b l e t o back p r e s s u r e .

t h e batance-measured l i f t and rake-measured t h r u s t d a t a as shown i n t h e f i g u r e s f o r t h e two l i f t / c r u i s e m i c s is a t t r i b u t e d solely t o n o d e l suckdown. As shown i n t h e two f i g u r e s , approximately 152 and 20% suckdown force w a s measared f o r : t h e two r e s p e c t i v e u n i t s a t t h e gear h e i g n t o f 3.3 f e e t ( H / D f 1.03). Tne nose MCDOUNELL AlRCRAPT COMPANY 9-35 MDCA4318 FIGURE 9-21 EFFECT OF GROUND HEIGHT ON INDIVIDUAL UNIT MEASURED LIFT Left LiWCruise Unit Only a = oo " 1200 1600 2000 2400 2800 3200 3600 4000 -. - ._ Corrected Fan Speed, N F / q - rpm GP70-0622 131 MCOONNELL AIRCRAFT COMPANV 9-26 MDC A431 8 FIGURE 9-22 EFFECT OF GROUND HEIGHT ON INDIVIDUAL UNIT MEASURED LIFT Right LiftICruise Unit Only lV=O0 1 200 - 1200 i600 2000 2400 2800 3200 3600 4000 Corrected Fan Speed, N F l n 0 - rpm O P 7 6 ~ 0 6 2 Z ~ l 3 2 MC00NNEL.L AtRCHAPT COMPANY 9 - 2 7 MDC A431 8 FIGURE 9-23 EFFECT OF GROUND HEIGHT ON INDIVIDUAL UNIT MEASURED LIFT Nose Lift Unit Only a = oo 1 600 a I

I

I , I

I .

1200 1600 2000 2400 2800 3200 3600 4000

-

Corrected Fan Speed, N F ; V U T ~ - rpm

GP78 0622 133 MDC A431 8 FIGURE 9-24 EFFECT OF GROUND HEIGHT ON INDIVIDUAL UNIT IDEAL THRUST Left LifdCruise Unit Only a=O0

Corrected Fan Speed, NF/% - rpm

MCDONNKLL AIUCRAFT COMPANY 9-29 MDCA4318 FIGURE 9-25 EFFECT OF GROUND HEIGHT ON INDIVIDUAL UNIT IDEAL THRUST Right LifdCruise Unit Only & S O 0 I200 1600 2000 2400 2800 3200 3600 4000 Corrected Fan Speed, N F l y ' B ; ; ; . rprn OP76.0622 138 MCOONNSLL AIRCRAFT COMPANV 9- 30 MDCA4318 FIGURE 9-26 EFFECT OF GROUND HEIGHT ON INDIVIDUAL UNIT IDEAL - - _ THRUST Nose L i f t Unit Only

._. - -

a = oo Corrected Fan Speed, N F / < - rpm OP76-0122.3 36 MCOONNELL AJRCIPAPT COMPANY 9-31 MDCA4318 FIGURE 9-27 EFFECT OF GROUND HEIGHT ON INDIVIDUAL UNIT LIFT AND THRUST Left LifdCruise Unit Only a 5 0 ’ NF/% = 3600 RPM

E

f -10

- 20

-30 Model Height Ratio. Ht’D GP79.0022 13T MCDONNELI AMCRAFT COMPANY 9-32 . . . .. - .._ .

MDCA4318

I

FIGURE 9-28 EFFECT OF GROUND HEIGHT ON INDIVIDUAL UNIT LIFT AND THRUST _ _ _ Right LifdCruise Unit Only Q 0 ' N F I K ~ 3600 RPM E X C .- 8 l C m

5 -10

- 20

-.30 0 1 2 3 4 5 6 7 Model Height Ratio - H!D I MCDONNELL AIIRCRAP T COMPANY 9-33 MDCA4318

FIGURE 9-29

EFFECT OF GROUND HEIGHT ON INDIVIDUAL UNIT LIFT AND THRUST

Nose Lift Unit Only Q E 0 ' NF/- 3600 RPM c X c .- 0 1 2 3 4 5 6 7 Model Height Ratio - H i 0 op7e-oezz. 7 39 MCDONNELI. AIRCRAFT COMPANY 9-34 MDC A4318 unit d a t a of F i g u r e 9-29 shows an e n t i r e l y d i f f e r e n t p i c t u r e . The r-ke-measured t h r u s t v a r i a t i o n is shown t o i n c r e a s e s t e a d i l y as t h e ground is approached. h i s i n c r e a s e is ca*!sed by movement t o a more f a v o r a b l e o p e r a t i n g p o i n t on t h e f a n map due t o p o s i t i v e back p r e s s u r e e f f e c t s . The n e g a t i v e d i f f e r e n c e s between t h e b a l a n c e l i f t and t h r u s t data at t h e i n t e r m e d i a t e H / D , l i k e t h e l i f t / c r u i s e u n i t s , are a t t r i b - uted t o suckd .. The p o s i t i v e d i f f e r e n c e between t h e l i f t and t h r u s t data a t t h e lowest H/D I similar t o t h a t experienced i n small s c a l e model f a n tests (Reference 4) a t low H/D. The e x p l a n a t i o n f o r t h i s phenomenon was tk.'.at as t h e h o r i z o n r a l l y mounted l i f t f a n approach t h e ground, a highe: than ambient b a s e p r e s s u r e w a s induced on t h e e x i t hub of t h e l i f t f a n , and a n e t p o s i t i v e l i f t was achieveu.

This same e 'ect is b e l i e v e d t o b e o c c u r r i n g on t h e nose u n i t on t h i s p a r t i c u l a r model c o n f i g u r a t i o n .

L i f t Loss summary A summary of t h e balance-measured l i f t d a t a (ALlL) shown on t h e previous summary p l o t s i s p r e s z n t e d i n Figure 9-30 f o r b o t h t h e t h r e e u n i t o p e r a t i o n d a t a and t h e s i n g l e u n i t l e f t l i f t / c r v i s e u n i t aata. The d a t a from t h i s t e s t are p l o t t e d on an e x i s t i n g comparison p l o t of previous m u l t i p l e - j e t model tests as obtained from Reference ( 6 ) . As shown i n t h e .':igure, tfie d a t a t r e n d s seen i n t h i s test are similar t o t h o s e shown f o r t h e previous tests. It should be noted, however, tiat t h e comparison d a t a from t h i s test are shown on t h i s e x i s t i n g summary p l o t f o r i l l u s t r a t i o n purposes o n l y , and do n o t n e c e s s a r i l y r e f l e c t t e P t d a t a acquired i n l i k e manner.

I I MCDONNELI AIRCRAFT COMPfiNY 9-35 FIGURE 9-30 LIFT LOSS IN GROUND EFFECT COMPARISONS Figure Reproduced from Reference (6) I I 1 i J

-1.0 -

0 2 4 6 8 13 HID CD74.6621 273 MCDOUNELL AJRCRAPT COMPANY 9- 35 9.3 GROLXD EFFECTS ON IXET REIXGESTION The e f f e c t s of ground height and s e l e c t e d m d e l t e s t v a r i a b l e s on i n l e t temperature reSnpestion c h a r a c t e r i s t i c s are p r e s e n t - 3 i n t h i s s e c t i o n f o r t h e f o u r inlets i n o p e r a t i o n on t h i s test model as shown i n Figure 9-31. The test v a r i a b l e s i n c l u d e t h e model ground h e i g h t , f a n speeds, ilose a n i t v e c t o r angle, yaw :-de s p l a y angle, inlet s h i e l d s i z e and s h i e l d d e f l e c t i o n angle. Xost of t h e d a t a p r e s e n t e d w a s measured u t i l i z i n g t h e Vidar d i g i t a l d a t a a c a u i s i t i o n system and is assumed to be steady state. S e l e c t e d analog d a t a traces are p r e s e n t e d f o r t h e more unsteady test runs and couparisons between t h e d i g i t a l and a n a l o g d a t a f o r t h e s e s e l e c t e d test runs are discussed.

Ground Height E f f e c t s The e f f e c t s of ground n e i g h t on i n l e t t t i p e r a t u r e r e i n g e s t i c n are p r e s e n t e d throughout t h i s s e c t i o n i n o r d e r t o p r o v i d e a b a s i s f o r comparing t h e e f f e c t s of t h e o t h e r t e s t v a r i a b l e s . A s p e c i f i c c o q a r i s o n of t h e i n l e t temperature rise f o r each of t h e f o u r inlets evaluated i n t h e test program is p r e s e n t e d as 2 f u n c t i o n of model ground h e i g h t r a t i o (H/D) i n F i g u r e 9-32. The data shown are presented f o r a c o n s t a n t ambient c o r r e c t e d f a n speed of 3600 rpm. The i n l e t temp- e r a t u r e rise index, ST,/AT is u t i l i z e & on t h i s and on every f i g u r e p r e s e n t e d j ' h e r e i n i n o r d e r t o ncn-dimensionalize t h e r e i n g e s t i o n d a t a for d i r e c t comparison purposes between u s . The temperature rise indeh removes from c o n s i d e r a t i o n t h e strons e f f e c t s of changes i n t h e exhatist j e t temperature, which w a s c o n t i n u a l l y changing over t h e course of t h i s L e s t program, due b o t h t o ambient t e m p e r a t u r e v a r i a t i o n s and t o t h e l a r g e jet temperature rise P s s o c i a t e d w i t h l a r g e i n l e t temperature rises. The 3Ti e q u a l s t . 2 i n l e t t o t a l temperature rise above ambient ), and t h e hT equals t h e t h r e e u n i t mass a u e r a p d j e t t e m p e r a t u r e

(TT2 - T a d J

rise above ambient (TT As shown i n F i g u r e 9-32, t h e peak i n l e t tempera- - Tamb ).

j t u r e rise occurred a t t h e i n t e m e d i a t e 8.3 f o o t h e i g h t t e s t e d f o r zll i n l e t s exceFt t h e l e f t l L f t / c r u i s e f a n i n l e t . The reasc;? for t h e lower i n l e t temperatures a t t h e 3.1 v e r s u s t h e 8.3 f o o t h e i g h t i s b e l i e v e d t o be t h e s t r o n g e r upwash dynamic p r e s s u r e and hence a s t r o n g e r l a t e r a l d e f l c c t i o n of t h e f l o w f i e l d by t h e a i r c r a f t .

Also shown i n t h e f i g u r e i s t h e i n l e t which s u s t a i n s t h e h i g h e s t r e i c g e s t i o n The approximate l o c a t i o n temperature rise, t h a t of t h e l e f t gas generator i n l e t .

of t h e l a t e r a l f o u n t a i n (see Figure 9-31) occurs, as expected, i n t h e v i c i n i t y of :.his i n l e t which is a l s o the only i n l e t without any f o m of a i r f r a m e s h i e l d i n g .

The D ~ L S S averag.. j e t temperatures a l o n g w i t h t h e ambiect temperatures f o r each d a t a p?l.?t cre presented i n a l l f i g u r e s i n t h i s s e c t i o n . I n general, t h e LT. v a l u e J MCDONNELL. AICICRAPT COMPANY 9-37 MDCA4318 FIGURE 3-31 POWERED MOOEL TEMPERATURE INSTRUMENTED INLETS Inlet Reingestion Tests Lateral Fountain 1-- Fountain Left Gas Generator I --- ' Gas Generator Inlet Left Lift/Cruise OP76-0622.275 MCDONNELL AIRCRAFT COMPANV 9- 38 MDCA4318 FIGURE 9-32 EFFE? C)F GROUND HEIGHT ON INLET REINGESTION NF/v% = 3600 RPM 0 1 2 3 4 5 6 7 Model Hetght Ratio, HID 5P76.0622.235 MCOONNhLL AIRCRAFT COMPANY 9-39 MDC A4318 measured i n most of t h e tests w a s approximately 100 Fahrenheit degrees.

Louver D e f l e c t i o n E f f e c t s on i n l e : r e i n g e s - The e f f e c t s of v a r y i n g nose u n i t louver d e f l e c t i o n angles t i o n is p r e s e n t e d f o r each i n l e t i n Figures 9-33, 9-34, and 9-35 for ground h e i g h t s of 3 . 3 , 8 . 3 , and 21.0 f e e t , r e s p e c t i v e l y . A l l d a t a are p r e s e n t e d f o r a c o n s t a n t c o r r e c t e d f a n speed of 3600 rpm. I n g e n e r a l for a l l i n l e t s a t a l l a l t i t u d e s , w i t h t h e e x c e p t i o n of t h e nose f a n i n l e t a t t h e 8.3 f o o t h e i g h t , d e f l e c t i n g t h e l o u v e r s from 80' t o 102" caused a s r e a d i l y i n c r e a s i n g and s i g n i f i - cant rise i n t h e i n l e t temperature l e v e l s msasured. These c h a r a c t e r i s t i c s were due t o t h e d i r e c t movement of t h e lateral fountain. Moving it rearward under t h e wing i n c r e a s e d s h i e l d i n g whereas moving it. forward away from i h e wing reduced s h i e l d i n g e f f e c t s . The s e n s i t i v i t y of inlet temperature rise t o louver d e f l e c t i o n f o r t h e more critical gas g e n e r a t o r i n l e t s w a s shown t o be much less as a l t i t u d e w a s increased.

D i f f e r e n t i a l Fan Speed E ' 's The e f f e c t of nosL sad v a r i a t i o n s on r e i n g e s t i o n by :.he f o u r i n l e t s is p r e s e n t e d i n Figures % - ,. 9-37, and 9-38 f o r ground h e i g h t s of 21.0, 5 . 3 and 3 . 5 f e e t , r e s p e c t i v e l y . The o s e fan speed was v a r i e d below and abovt a nominal The e f f e c t s were found t o b e most pronounced a t t h e 3 . 3 f a n speed of 3600 rpm.

f o o t ground height. Decreasing t h e nose f a n speed and hence its j e t dynamic p r e s s u r e caused t h e lateral f o u n t a i n t o s h i f t forward from beneath t h e s h i e l d i n g wing area and t h e r e f o r e i n c r e a s e t h e r e i n g e s t i o n l e v e l s of a l l f o u r inlets.

I n c r e a s i n g t h e nose f a n speed moved t h e f o u n t a i n rearward and decreased r e i n g e s - t i o n . S i m i l a r c h a r a c t e r i s t i c s were observed a t tl-3 8.3 f o o t ground h e i g h t , except t h a t a d e c r e a s e i n f a n speed below t h e nominal 3600 rpm v a l u e showed very l i t t l e e f f e c t . Both t h e nose and l i f t / c r u i s e f a n speed v a r i a t i o n e f f e c t s a t t h e 21.0 f o p t h e i g h t , Figure 9-36, showed only minor e f f e c t s on i n l e t r e i n g e s - tim.

Nozzle Exhaust Splaying E f f e m The e f f e c t s of s p l a y i n g t h e nozzle j e t exhausts on i n l e t r e i n g e s t i o n temperature The d a t a are p r e s e n t e d i n Figures 9-39 through 9-42 f o r t h e four i n l e t s t e s t e d .

a r e p r e s e n t e d versus (H/D) f o r a c o n s t a n t c o r r e c t e d f a n t h e ground h e i g h t r a t i o speed of 3600 rpm.

Splaying was achieved using t h e manually positioned yaw vanes on each l i f t The two yaw vanes on each l i f t c r u i s e u n i t and the two on t h e nose u n i t u n i t .

splayed outboard 1 2 ' .

MCOONNELL AIRCRAP T COMPA N V 9-40 MDC A43 1 8 FIGURE 9-33 EFFECT OF NOSE UNIT DEFLECTION ANGLE ON INLET REINGESTION lvlodei tieight = 21.0 F?

-. . ~-- _-.

0 . 3 0 . 2 0.1 0 1 2 3 4 5 6 7 Model Height R, 0, H:D CP76 5621 1 4 3 MCDONNELL A m c H a F r COMPANY 9-47 MDC A431 8 FIGURE 9-40 EFFECT OF NOZZLE EXHAUST SPLAYING ON INLET REINGESTION Forward Gas Generator Inlet -. ..

NF’IL ?(T, 3600 SPhl 0 . 9 0.8 0.7 0 . 6 L - 0 . 3 0.2 0.1 0 1 2 3 4 5 6 Model Heiyht R a t i 3 , H,D OP76 0622 7 1 2 MCDONNELL AIRCHA F f COMPANY

+$a

MDC A431 8 FIGURE 9 4 1 EFFECT OF N O Z Z L E EXHAUST SPLAYING ON INLET REINGESTION L/H Lift/Cruise F a n i n l e t - NF/\ "T, = 3600 RPM Yaw Vane Deflections 'iiew Looking Forward Left Lift/ Noye Liii Right Lift Cruise Unit 'Jnit Cruise U n i t - 0 1 2 3 4 5 6 7 %lode1 Heibnt Eatio, ti 0 GP76 0 6 2 2 2 4 5 MCDCNNELL AIRCRAFT C O M C ' 4 f d Y 9-49 MDC A431 8 I FIGURE 9*Q?

EFFECT OC NOZZLE EXHALIST SPLAY IPlG ON I N L E T R E I N G E S T I 9 N Nose Fan Inlet N F ; \ " T = 3600 RP'k?

MK: A4318 As shown I n t h e f i g u r e s f o r each i n l e t , t h e e f f e c t of s p l a y i n g v a s v e r y I b z n e f i c i a l i n reducing r e i n g e s t i o n a t t h e two lover ground h e i g h t s f o r a l l i n l e t s .

A t t h e 21.0 f o o t h e i g h t , however, s p l a y i n g was shown t o have a n e g a t i v e e f f e c t i n t h a t t h e temperature rise on each i n l e t w a s increased. T h i s is b e l i e v e d to b e due t o inducing a s t r o n g e r far f i e l d r e c i r c u l a t i o n f l o w f i e l d around t h e model a t t h e h i g h e r a l t i t u d e s .

E f f e c t s of I n l e t P h i t r .-s

The e f f e c t s of t h e small gas g e n e r a t o r i n l e t s h i e l d on r e i n g e s t i o n temperature for tests at t h e 8.3 f o o t ground h e i g h t are p r e s e n t e d f o r t h e two gas g e n e r a t o r ialc'.^s i n F i g u r e 9-43 znd f o r t h e two f - n i n l e t s i n F i g u r e 9-44.

The e f f e c t s of b i - n l a r g e and small gas g e n e r a t o r i n l e t s h i e l d i n g on i n l e t r e i n g e s t i o n temperature f o r tests a t t h e 3.3 f o o t ground h e i g h t are p r e q t e d f o r t h e two gas g e n e r a t o r s i n l e t s i n Figure 9-45 and f u r t h e tu0 f a n i n l e t s i n Figure 9-46. Both s h i e l d s were found : o b e v e r y e f f e c t i v e i n reducing t h e l e f t gas generator f n l e t r e i n g e s t i o n at t h e 3.3 f o o t ground h e i g h t , as shown i n t h e Tn Figure 9-46 t h e gas g e n e r a t o r f n l e t s h i e l d s were f o m d t o have essen- f i g u r e .

t i a l l y nc e f f e c t on t h e r e i n g e s t i o n c h a r a c t e r i s t i c s of b u t h t h e l i f t / c r u i s e and nose fan i n l e t s . A photo of t h e l a r g e s h i e l d i n s t a l l e d on t h e powered model at t h e 3.3 f o o t h e i g h t is shown i n F i g u r e 5-5.

(ss) on i n l e t r e i n g e s t i o n are p r e c m t e d The effects of s h i e l d d e f l e c t i o n a n g l e in F i g u r e 9-47 a t a f a n speed of 2500 rpm at a ground h e i g h t of 3.3 f e e t . S h i e l d d e f l e c t i o n ang?e tests were conducted w i t h t h e l a r g e s h i e l d only a t t h i s lower ground height. As shown i n t h e f i g u r e , a s h i e l d d e f l e c t i o n a n g l e of 0" vas found t o provide t h e b e s t s h i e l d i n g f o r b o t h gas g e n e r a t o r i n l e t s . The e f f e c t s of s h i e l d d e f l e c t i o n angle on bct' fan i n l e t s were found t o b e minor.

Analog Temperature Data A r e p r e s e n t a t i v e s e l e c t i o n of t h z c o n t i n u o u s l y recorded analog temperature d a t a measured d u r i n g t h e s t a t i c test program is p r e s e n t e d i n t h i s s e c t i o n . i n d i v i - d u a l temperature traces are p r e s e n t e d f a r s e l e c t e d thermocouples l o c a t e d i n each of the f o u r i n l e t s i n the r e i n g e s t i o n tests. Figure 9-48 shows t h e i n l e t t h e n o c o u p l e numbering used f o r i d e n t f f y i n g the analog t r a c e s p r e s e n t e d h e r e i n .

The purpose of ? r e s e n t i n g t h e s e l e c t e d a n a l o g d a t a r - a c e s is t o i l l u s t r a t e . - the i n l a t temperature v a r i a t i o n d u r i n g an i n d i v i d u a l test p o i n t , and a l s o t o show t h e d e v i a t i o c from the s i n g l e scan Vidar systzm measured i n l e t temperatures. For a l l analog d a t a presented, the v a r i a t i o n s are shovn duri1.g the complete 90 second time d u r a t i o n of t h e Vidar d i g i t a l d a t a a c q u i s i t i o n process. During each t e s t p o i n t , the model p r e s s u r e s , s e l e c t e d temperatures, f a n speeds, and load c e l l d a t a MCOONUSLL AIRCRAP COMPANY 9-51 MDC A4318 are a l l measured d u r i n g t h e i n i t i a l 85 seconds of t h e test p o i n t . The last 5 seconds a r e used t o measure t h e i n l e t temperatures d i g i t a l l y .

The i n l e t r e i n g e s t i o n d a t a of t h i s r e p o r t are a l l based on t h e s e d i g i t a l p r e s e n t e d i n t h e previous f i g u r e s measurements.

traces f o r some thermocouple from each of the f o u r i n l e t s t e s t e d Analog d a t a are presei.ted i n Figures 9-49, 9-50, and 9-51 f o r t h e t h r e e model ground h e i g h t s t e s t e d of 21.0, 8 . 3 , and 3.3 f e e t . The temperature v a r i a t i o n s are shown relative to ths d i g i t a l l y measured r e f e r e n c e temperature ( T m ) f o r each thermocouple.

The t i m e at which t h e d i g i t a l TREF is measured is a l s o i n d i c a t e d i n each f i g u r e f o r each thermocouple. The p a r t i c u l a r t e w e r a t u r e s s e l e c t e d and shown h e r e are t h e "worst ease" v a r i a t i o n s f o r each i n l e t a t t h e s e test c o n d i t i o n s . As shown, t h e runs and also t h e d e v i a t i o n s from t h e temperature v a r i a t i o n s d u r i n g t h e t h r e e were both found t o be w i t h i n 2 10OF.

d i g i t a l l y measured temperature l e v e l s The analog temperature d a t a f o r a l l '.hemnocouples i n s t a l l e d i n t h e a o r e c r i t i c a l l e f t hand gas g e n e r a t o r i n l e t are p r e s e n t e d i n F i g u r e s 9-52, 9-53, and 9-54 f o r t h e t h r e e model ground h e i g h t s of 21.0, 8.3, and 3.3 f e e t . Thermocouple No. 6 on t h e L/H gas g e n e l a t o r i n l e t w a s i n t e r n i t t a n t l y i n o p e r a t i v e d x r i n q t h e s t a t i c tests, and t h e r e f o r e is noc shown. ?or t h e seven t h e n o c o u p l e n e a s u r e z e n t s shown i n t h e s e t h r e e f i g u r e s , t h e temperature v a r i a t i o n s during a given r u n were found to be reasonably s t a b l e , and t h e d i g i t a l l y measured Vidar i n l e t t e u p e r a t u r e s are a l s o sliovin t o be very r e p r e s e n t a t i v e measurements or' t h e l e v e l of r e i n g e s t i o n o c c u r r i n g w i t h t h i s test model.

9 . 4 FLOW VISUALIZATIOX TESTS Flow v i s u a l i z a t i o n tests were conducted a t t h e 21.0 and 3.3 f o o t model h e i g h t s , u t i l i z i n g CONUS o i l for smoke generation. F l o w f i e l d p a t t e r n s were recorded w i t h h i g h speed (250 t o 500 frames/sec) and normal speed (24 f r a u e s / s e c ) movie cameras.

S t i l l photos were a l s o taken a t s e l e c t e d c o n d i t i o n s . F i g u r e 9-48 shows t y p i c a l still photo comparisons of t h e f l o w f i e l d i n t h e v i c i n i t y of t h e L/H gas g e n e r a t o r i n l e t w i t h and w i t h o u t i n l e t s h i e l d s i n s t a l l e d . Observations made during those tests c l e a r l y show when ar,d where i n l e t r e i n g e s t i o n i s o c c u r r i n g and l i k e w i s e support t h e i n l e t r e i n g e s t i o n trends p r a s e n t e d i n t h i s r e p o r t .

- - ._

. - _ .

MCOONNELI AIRCRAF I COMPANV 9-51.4 MDCA4318 FIGUR 5 - 9 - 4 3 EFFECT OF SHIELDING ON INLET REINGESTION : Mode! Height = 8.3 h 1600 2000 7400 2800 3200 3600 4000 4200

Corrected Fan Speed, N F ; ~ - rpm

OC76 0622.248 MDCA4318 FIGURE 9-44 EFFECT OF WIELDING ON INLET REINGESTION Model Height = 8 . 3 h lllose Fan Inlet 0.4 0.3 0.2 0.1 k

%

p 1600 2400 a800 3200 3600 4000 4200

E c

<

c

-. .

MCDONNKLL a m c R ~ F r COMPANY 9-58 MDC A431 8

f

MCVOUNSAL AIRCRAFT COMPANY 9-59 MOC A431 8

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+ I + I MCDONNLLL AIRCRAFT COMPANY 9-60 MDC A431 8 hl o A e3

v

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MCDONNlLL AIRCNAFT COMPANV 9-61 MDC A4318 m (6 I

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I !

I I

I I u .

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MDC A4318 m M P r"

-x

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w WJ -0 c , I i 9-63 HDC A4318 10. CONCiUSIONS The complete three-fan powered l i f t confi.guration e x h i b i t s p o s i t i v e induced I .

l i f t (AL/FG from 0 t o 0.1) f o r o p e r a t i o n a l combinations of t h r u s t v e c t o r angle and j e t v e l o c i t y r a t i o (V0/Vj). The corresponding induced d r a g is g e n e r a l l y negative.

2. The two-fan powered l i f t c o n f i g u r a t i o n w i t h t h e h o r i z o n t a l t a i l o f f and t h e nose l i f t u n i t s e a l e d e x h i b i t s l a r g e p o s i t i v e induced l i f t values. With t h e l i f t / c r u i s e unit a t go", AL/FG v a r i e s from a nominal 0.1 t o 1.7 a t j e t v e l o c i t y r a t i o s o f 0.1 t o 0.7.

3. Operation of t h e nose l i f t unit r e s u l t s i n a s i g n i f i c a n t adverse e f f e c t QU induced l i f t .

4 . A comparison of l i f t and d r a g d a t a from s t a t i c tests of three-fan coubined o p e r a t i o n and t h r e e i n d i v i d u a l f a n o p e r a t i o n s i n d i c a t e s d i f f e r e n c e s i n from a mutual i n t e r z c t i o n betweeil induced aerodynamic f o r c e s t h a t may r e s u l t D i f f e r e n c e s i n l i f t range from -5Z t o +2X o f t o t a l t h e p r o p l s i o n units.

t h r u s t and d i f z e r e n c e s i n d r a g range from -62 t o -3Z 0 5 t o t a l t h r u s t .

5 . The three-fan potqered l i f t c o n f i g u r a t i o n has u n s t a b l e d i r e c t i o n a l s t a b i l i t y at very low Vo/VJ r a t i o s due t o t h e d e s t a b i l i z i n g e f f e c t of t h e nose l i f t A t h i g h e r jet v e l o c i t y r a t i o s , t h e c o n f i g u r a t i o n is s t a b l e u n i t ran drag.

both d i r e c t i o n a l l y and l a t e r a l l y .

6 . The aerodynamic l i f t c o n f i g u r a t i o n w i t h h o r i z o n t a l t a i l on h a s s t a t i c l o n g l t u d i n a l s t a b i l i z y 8t angles of a t t a c k up t o wing-stall a n g l e of a t t a c k .

Horizontal t a i l c o n t r o l e f f e c t i v e n e s s i s r e t a i n e d through t h e h i g h e s t a n g l e of a t t a c k t e s t e d .

7. The aerodyaamic l i f t c o n f i g u r a t i o n ha5 n e u t r a l lateral s t a b i l i t y and s t a b l e Above 0 ' a n g l e of a t t a c k t h e d i r e c t i o n d s t a b i l i t y a t 0 ' a n g l e of a t t a c k .

c o n f i g u r a t i o n is s t a b l e both d i r e c t i o n a l l y and l a t e r a l l y through t h e h i g h e s t angle of a t t a c k t e s t e d .

The e f f e c t of ground h e i g h t on t o t a l l i f t loss f o r t h e modal t e s t e d vas less 8.

than 1% while o p e r a t i n g a t constant fan speeds.

I n l e t r e i n g e s t i o n 1 .rels tiere s t r o n g l y a f f e - t e d by t h e model v a r i a b l e s t e s t e d 9.

i n t h i s program, i n c l u d i n g model h e i z n t , nos2 v e c & o r an.gle, d i f f e r e n t i a l f a n speeds, s p l a y i n g , and i n l e t s h i e l d i n g .

g e n e r a t o r i n l e t s h i e l d s appear t o be effective devices for reducing L 10. Gas B i n l e t r e i n g e s t i o n l e v e l s a t low ground h e i & t s .

MCDONNelS AIRCRAFT COMPANY 10- 1 HDC A4318 1 1 . The l i f t / c r u i s e fan, nose fan, and gas generator i n l e t s were a l l found to exhibit high levels of inlet recovery coupled with low distortioa throughout the range of tests performed on t h i s model, particularly a t angles of attack less than 20°.

MCDONNELL AlRCRAFY COMPANY 10-2 FfDC A4318 11. LIST OF REFERENCES 1. B . J. Gambucci, K. Aoyagi, L. S . Rolls, "Wind Tunnel I n v e s t i g a t i o n of a Large Scale Hodel o f a Lif t / C r u i s e Fan V/STOL A i r c r a f t , " NASA n S - 7 3 , 139 d a t e d Yay 1976.

"Guide f o r Pla-uring I n v e s t i g a t i o n s in t h e Ames 40- By 80-Foot Wind Tunnel," 2.

dated June 1975.

J. K . ' O r r , C. W . Sapp, %ode1 260 Low Speed Wind Tunnel Aerodynamic T e s t Data 3.

A S U Configuration," Report l4DC B9827-2, dated December 1975.

E . J. Phillips, "Model 260 L i f t / C r u i s e I n l e t S t a t i c and High Speed Performance 4 .

T e s t - S e r i e s I," R e p o r t MDC B9856-2, d a t e d December 1974.

"LF460 Decaf1 Design - F i a n l Technical Report on Aircraft Support A c t i v i t y , " 5.

Report NASA CR-121146, General Electric Co., dated A p r i l 1973.

6. Richard J. Margason, "Review of Propulsion-Induced E f f e c t s on Aerodynamics of Jet/STOL A i r c r a f t , " NASA TX D-5617, d2" ' Cebruary 1970.

.-_ . . .

MCDONNlLL AJRCHAPT c 3 M P A I Y Y 11- 1

Next: APPENDIX A →

APPENDIX A

MDC A4318

WIND TUNNEL AND GROUND STATIC INVESTIGATION

OF A LARGE SCXLE-MODEL OF A LIFT/CRUISE FA3 V/STOL AIRCB,MT

APPENDIX A

40’ X 80’ WIND TUNNEL

TEST SCHEDULE AXD BALANCE DATA

I MCDOUUCLL AlRCRAFT COMPAUk ,

A- 1

NDC A4318

EXPLAMATlON OF SYMBOLS FOR RUN SCHEDULE

Symbol other RUN PT Point Number

ALPHA Angle of Attad;: a -4O te 320

BET#$ Sideslip Angle. 8 -4°toaoo

b , DLC . ,ihKruise Unit Geometric

Deflection. 8 LC oOmsoO

* BNL ~ o s e ~ i t t Unit +metric

Deflection. S N L oo to toso

DF Flap Deflection, S f oOt0300 Left Ailwon Deflection, 8;, DAL - 2 9 to 25O Right Aileron Deflection, Sa DAR -25O to 25'

OH Horizontai Tail Deflection, 6 H -2OO to 20° I H = -99. Indicates that the horizontal

ail is removed and the flow survey rake nstalled

IH = 99. Indicates that the horizontal

ail is removed and the flow survey rake s also removed DR Rudder Deflection, 6 R -200 to 3 0 ' (Nominal) Corrected Test Section QC Dynamic Pressure, q, 0 to 49.2 psf Nose Lift Unit Corrected FAN 1 Fan Speed, N ~ / f i ~ ~

0 to 3600 rpm

FAN 2 Lft Lift/Cruise Unit

Corrected Fan speed, NF/ ,/T 0 to 4100 rprn

T, Ri&t Lift/Cruise Unit FAN 3

Corrected f a r Speed, NF/ .J?$ 0 to 4100 rpm

c ._.

DYLC LiftICruise Unit Yaw Vane )YLC = 99. Indicates lift cruise unit yav Deflection, 6yLc -la0 to 120 ana removed

DYNL Nose Lift Unit Yaw Vane

)YNL = 99. Indicates nose lift unit yaw Deflection, 6yNL - 1 2 O to la0 anes removed.

.- Code for Nose Gear NOSE

GEAR 1 - Nose Gear Off

2 - Nose Gear On

NOSE U N l l Code for Nose Unit Inlet Covers

INLET 1 - Nose Unit lnlets Cover ?d

COVERS 2 - Nose Unit Inlets Open

I#CDOU&ELL AIRCRAC t COMPANV

A- 2

IbYWL FLbl PAM3 OYLC m L I)o?l HCR DE6 O t 6 WB ' 0 . 0 -00 0. 11.

e. B.00 ore 0.

34. e.0 o g r e a. 8-80 0.

99.0 43. Q.0 6. 0.08 0.

OQ.0 8.0 0.

8: 0.0 09.0 e. a : % ' 2 20%: e.

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@.eo 0. 291 1. 0.8 0.

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4 i 3 8.8

9. ,?. 3s 3641. 0.6 eQ.0 4 9.0 0 .

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n.n -n.o p a 0.0

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0.9 1 -54.

7 15 -9 . 3 0 . 2 2

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Next: APPENDIX B →

APPENDIX B

M)C A4318 WIND m L AXD GROUND STATIC INVESTIGATION OF A LARGE SCALE-MODEL OF A LIETICRUISE FAN V/STOL AIRCRAFT APPENDIX B OUTSIDE STATIC TEST RUN SUMMARY MCOONNIILL LIlRCRAPT COMPANY 13- I . .

B-2 .

MDC A431 8 ' B-4 MDCA4318

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MDCA4318

ORIGINAL PAGE I S

OF POOR QUALID

MDCA4318

Overview

Section View in PDF

SYMBOLS AYD NOMENCLATURE . . . . . . . . . . . . . . . . . . . xv

IX'RODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . 1-1

1 . 1 Background . . . . . . . . . . . . . . . . . . . . . . . . 1-1

1 . 2 40' x 8 0 ' Wind Tunnel Tests . . . . . . . . . . . . . . . 1-2

1 . 3 Outside Static Tests . . . . . . . . . . . . . . . . . . . 1-2

OUTSIDE STATIC TEST FACILITY . . . . . . . . . . . . . . . . . . 5-1

5 . 2 Data Acquisition System . . . . . . . . . . . . . . . . . 5-7

WIND TUNIEL TEST P R O G W Y . . . . . . . . . . . . . . . . . . . 6-1

6.1 Test Conditions . . . . . . . . . . . . . . . . . . . . . 6-1

6.2 Test Procedures . . . . . . . . . . . . . . . . . . . . . 6-4

6.3 Data Reduction . . . . . . . . . . . . . . . . . . . . . . 6-6

8.3 Powered Lift Configuration - Induced Lift and Drag

8.4 Powered L i f t Configuration . Pitching Yompnt

8.5 Powered Lift Configuration . Lateral-Directional

Characteristics . . . . . . . . . . . . . . . . . . . . 8-126

-

8.6 Aerodynamic L i f t Configuration - L o n g i t u d i n a l

C h a r a c t e r i s t i c s . . '. . . . . . . . . . . . . . . . . . 8-152

8.7 Aerodynamic L i f t Configuration - L a t e r a l - D i r e c t i o n a l

C h a r a c t e r i s t i c s . . . . . . . . . . . . . . . . . . . . 8-173

8.8 A i r I n d u c t i o n System Performance. . . . . . . . . . . . . 8-191

9. OUTSIDE STATIC TEST RESULTS. . . . . . . . . . . . . . . . . . 9-1

9 . 1 Propulsion System C a l i b r a t i o n s . . . . . . . . . . . . . . 9-1

9 . 2 Ground E f f e c t s on A i r c r a f t L i f t Loss. . . . . . . . . . . 9-13

9.3 Grc;md E f f e c t s on I n l e t Reingestion . . . . . . . . . . . 9-37

9.4 Flow V i s u a l i z a t i o n Tests. . . . . . . . . . . . . . . . . 9-51A

10. CONCLUSIONS. . . . . . . . . . . . . . . . . . . . . . . . . . 10-1

11. LIST OF REFEREXCES . . . . . . . . . . . . . . . . . . . . . . 11-1

APPENDIX h 40' x 80' WIXD TUNNEL TEST SCHEDULE AXD BALMCE DATA . . . . . A- 1

APPENDIX B OUTSIDE STATIC TEST RW S U I W S . . . . . . . . . . . . . . . . B- 1

Large S c a l e Powered Nodel P r o p u l s i o n System . . . . . . . . . . . . 2-2

Large S c a l e Powered Node1 . . . . . . . . . . . . . . . . . . . . . 2- 3

L i f t / C r u i s e Fan I n l e t Design Geometry . . . . . . . . . . . . . . . 2- 5 2-8

. . . . . . . . . . 2-9

Gas Generator I n l e t S h i e l d Geometry . . . . . . . . . . . . . . . . 2- 7

Nose Fan I n l e t Design Geometry . . . . . . . . . . . . . . . . . . 2-8

Forward Engine I n l e t Design Geometry . . . . . . . . . . . . . . . 2-9

L i f t / C r u i s e Unit Vectoring System Geometry . . . . . . . . . . . . 2- 11 2-15

Nose L i f t Unit Vectoring System Geometry . . . . . . . . . . . . . 2-12

3-2 Location of Forward Fuselage S t a t i c P r e s s u r e P o r t s . . . . . . . . 3 3

3-9 Engine Face I n s t r u m e n t a t i o n . . . . . . . . . . . . . . . . . . . . 3-12

4- 1 40 x 80 Wind Tunnel Test F a c i l i t y . . . . . . . . . . . . . . . . . 4-2

Wind Tunnel . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-3

Wind Tunnel . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-4

Wind Tunnel . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-5

Wind Tunnel . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-6

Wind Tunnel . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-7

4- 7 40 x 80 Control Room Layout . . . . . . . . . . . . . . . . . . . . 4-8

5-2 Node1 Ground Height V a r i a t i o n s . . . . . . . . . . . . . . . . . . 5-3

Test F a c i l i t y . . . . . . . . . . . . . . . . . . . . . . . . . . 5-4

Test F a c i l i t y . . . . . . . . . . . . . . . . . . . . . . . . . . 5-5

System Components . . . . . . . . . . . . . . . . . . . . . . . . 6-9

8.1-1 Left Lift/Cruise Unit Static Thrust . . . . . . . . . . . . . . . . 8-3

8.1-2 Right Lift/Cruise Unit S t a t i c Thrust . . . . . . . . . . . . . . . 8-4

8.1-3 Nose L i f t Unit S t a t i c Thrust . . . . . . . . . . . . . . . . . . . 8-5

8.1-7 Nose L i f t Unit I d e a l Thrust . . . . . . . . . . . . . . . . . . . . 8-9

8.1-8 Nose L i f t Unit I d e a l Thrust . . . . . . . . . . . . . . . . . . . . 8-10

8.1-9 Thrust Calibration Ccefficient Determination . . . . . . . . . . . 8-11

8.1-10 L i f t l C r u i s e Unit Thrust Coefficients . . . . . . . . . . . . . . . 8-12

8.1-12 ',eft Lift/Cruise Unit Thrust Vector Angles . . . . . . . . . . . . 8-14

Nose L i f t Unit Thrust Vector Angles . . . . . . . . . . . . . . . . 8-18

Forward Speed Effect on LiftICruise Unit Performance . . . . . . . 8-27

a . 2-6

Forward Speed Effects on Various Nose Unit Vector Angles . . . . . 8-28

Mass Flow Variations with Forward Speed . . . . . . . . . . . . . . 8-31

5 m = 43'. SJ 2 0 . 1 ' . . . . . . . . . . . . . . . . . . . . . . 8-38

6 ~ t 0 43'. 8 j = 2 0 . 1 ' . . . . . . . . . . . . . . . . . . . . . . 8-39

5 m 43'. 85 ID 2 0 . 1 ' . . . . . . . . . . . . . . . . . . . . . . 8-40

6NL 43'. a j 2 0 . 1 ' . . . . 8-41

-

6 ~ ~ 9 0 ' . 8 ~ ~ 8 4 . 7 ' . . . 6 . . . . 8-67

6~ a O', 6 ~ c 23', 6 a 43', 8j = 20.1' . . . . . . . . . . . . 8-68

8H o', 6 ~ c m 38", 6m 43', 8J 29.2O . . . . . . . . . . . . 8-69

6g O', 6 ~ c = 56', 6 m 43', 85 44.5" . . . . . . . . . . . . 8-70

a 0'9 6Lc 23'9 6 m 43', 8J 20.1' . . 8-74

a O', 6Lc a 23', 6 m = 43', BJ = 20.1' . . . . . . . . . . . . 8-75

a 8'. 6 ~ c 23', 6 m 43', BJ p 20.1' . . . . . . . . . . . . 8-76

u 8' d ~ c 0 23', 6m 43', 85 20.1' . . 8-77

Section