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AIRPLANE DESIGN E=SPIOS=PPPIP=P

PART V: COMPONENT WEIGHT ESTIMATION PIPP3PPPPIPPDPIIIIPPPPI~~~~PIPPPaPPP

Dr. Jan Roskam Ackers Distinguished Professor

of Aerospace Engineering The University of Kansas

Lawrence, Kansas

NO PART OF THIS BOOK MAY BE REPRODUCED WITHOUT PERMISSION FROM THE AUTHOR

Copyright: Roskam Aviation and Engineering Corporation Rt4, Box 274, Ottawa, Kansas, 66067

Tel. 913-2421624 F i r s t Print ing: 1 9 8 5

TABLE OF CONTENTS IIPIPPIIPPIPPISPP

TABLE OF SYMBOLS v i i

ACKNOWLEDGEMENT x i i i

1. INTRODUCTION 1

2. CLASS I METHOD FOR ESTIMATING AIRPLANE COMPONENT WEIGHTS 3 2.1 A METHOD FOR ESTIMATING AIRPLANE COMPONENT

WEIGHTS WITH WEIGHT FRACTIONS 4 2.2 EXAMPLE APPLICATIONS 7

2.2.1 Twin Engine P r o p e l l e r Driven Ai rp lane 7 2.2.2 Je t Transpor t 10 2.2.3 F i g h t e r 12

3. CLASS I METHOD FOR ESTIMATING AIRPLANE INERTIAS 17 3.1 ESTIMATING MOMENTS OF INERTIA WITH RADII

OF GYRATION 1 7 3.2 EXAMPLE APPLICATIONS 19

3.2.1 Twin Engine P r o p e l l e r Driven Ai rp lane 19 3.2.2 Jet Transpor t 20 3.2.3 F i g h t e r 21

4. CLASS I1 METHOD FOR ESTIMATING AIRPLANE COMPONENT WEIGHTS 4.1 A METHOD FOR ESTIMATING AIRPLANE COMPONENT

WEIGHTS WITH WEIGHT EQUATIONS 4.2 METHODS FOR CONSTRUCTING V-n DIAGRAMS

4.2.1 V-n Diagram f o r FAR 23 C e r t i f i e d Airplanes 4.2.1.1 Determinat ion o f +lg s t a l l

speed, V 4.2.1.2 ~ e t e r m i n a t i o n of des ign

c r u i s i n g speed, V 4.2.1.3 Deterlllination of 6es ign

d iv ing speed, V 4.2.1.4 Determinat ion oP des ign

maneuvering speed, V 4.2.1.5 Determinat ion of negPt ive

s t a l l speed l i n e 4.2.1.6 Determinat ion of des ign

l i m i t l oad f a c t o r , nlim

4.2.1.7 Cons t ruc t ion of g u s t load f a c t o r l i n e s i n Fig.4.1 f 34

4.2.2 V-n Diagram f o r FAR 25 C e r t i f i e d Ai rp l anes 3 5

P a r t V Conten ts Page i

4.2.2.1 Determination of +lg stall speed. VSl 35

4.2.2.2 Determination of design c r u i s i n g speed. 3 5

4.2.2.3 Determination o f V ~ e s i g n diving speed, V 3 7

4.2.2.4 Determination OF design maneuvering speed, V 3 7

4.2.2.5 Determination of desfgn speed f o r maximum g u s t i n t e n s i t y . Vg 3 7

4.2.2.6 Determination of negat ive s t a l l speed l i n e 3 7

4.2.2.7 Determination of design l i m i t load f a c t o r . nlim 3 7

4.2.2.8 Construct ion of g u s t load f a c t o r l i n e s i n Fig.4.2b

4.2.3 V-n Diagram f o r M i l i t a r y Airplanes 4.2.4 Example Applicat ion

4.2.4.1 Twin Engine Prope l l e r Driven Airplane

4.2.4.2 Jet Transport 4.2.4.3 Fighter

4.3 EXAMPLE APPLICATIONS FOR CLASS I1 WEIGHT ESTIMATES 4.3.1 Twin Engine Prope l l e r Driven Airplane 4.3.2 J e t Transport 4.3.3 Fighter

5. CLASS I1 METHOD FOR ESTIMATING STRUCTURE WEIGHT 5.1 WING WEIGET ESTIMATION

5.1.1 General Aviation Airplanes 5.1.1.1 Cessna Method 5.1.1.2 USAF Method 5.1.1.3 Torenbeek Method

5.1.2 Commercial Transport Airplanes 5.1.2.1 GD Method 5.1.2.2 Torenbeek Method

5.1.3 M i l i t a r y P a t r o l , Bomb and Transport Airplanes

5.1.4 Fighter and Attack Airplanes 5.1.4.1 GD Method

5.2 EMPENNAGE WEIGHT ESTIMATION 5.2.1 General Aviation Airplanes

5.2.1.1 Cessna Method 5.2.1.2 USAF Method 5.2.1.3 Torenbeek Method

5.2.2 Commercial Transport Airplanes 5.2.2.1 GD Method 5.2.2.2 Torenbeek Method

P a r t V Convents Page ii

5 .2 .3 Mil i t a ry P a t r o l , Bomb and Transport Airplanes

5 . 2 . 4 Figh te r and Attack Airplanes 5 . 3 FUSELAGE WEIGHT ESTIMATION

5 . 3 . 1 General Aviation Airplanes 5 . 3 . 1 . 1 Cessna Method 5 . 3 . 1 . 2 USAF Method

5 . 3 . 2 Commercial Transport Airplanes 5 . 3 . 2 . 1 GD Method 5 . 3 . 2 . 2 Torenbeek Method

5 .3 .3 Mil i t a ry P a t r o l , Bomb and Transport Airplanes 5 . 3 . 3 . 1 GD Method

5.3 .4 Fighter and A t t a c k Airplanes 5.4 NACELLE WEIGHT ESTIMATION

5 . 4 . 1 General Aviation Airplanes 5 . 4 . 1 . 1 Cessna Method 5 . 4 . 1 . 2 USAF Method 5 .4 .1 .3 Torenbeek Method

5 . 4 . 2 Commercial Transport Airplanes 5 . 4 . 2 . 1 GD Method 5 . 4 . 2 . 2 Torenbeek Method

5 .4 .3 Mil i t a ry P a t r o l , Bomb and Transport Airplanes

5 .4 .4 Figh te r and A t t a c k Airplanes 5 . 5 LANDING GEAR WEIGHT ESTIMATION

5 . 5 . 1 General Aviation Airplanes 5 . 5 . 1 . 1 Cessna Method 5 . 5 . 1 . 2 USAF Method

5 . 5 . 2 Commercial Transport Airplanes 5 . 5 . 2 . 1 GD Method 5 .5 .2 .2 Torenbeek Method

5 .5 .3 M i l i t a r y P a t r o l , Bomb and Transport Airplanes

5 . 5 . 4 Figh te r and A t t a c k Airplanes

6. CLASS I1 METHOD FOR ESTIMATING POWERPLANT WEIGHT 6 . 1 ENGINE WEIGHT ESTIMATION

6 . 1 . 1 General Aviat ion Airplanes 6 . 1 . 1 . 1 Cessna Method 6 . 1 . 1 . 2 USAF Method 6 . 1 . 1 . 3 Torenbeek Method

6 .1 .2 Commercial Transport Airplanes 6 .1 .3 M i l i t a r y P a t r o l , Bomb and Transport

A i rp lanes 6 .1 .4 Figh te r and A t t a c k Airplanes se

6.2 A I R INDUCTION SYSTEM WEIGHT ESTIMATION -

6 . 2 . 1 General Aviation Airplanes 6 . 2 . 1 . 1 Cessna Method 6 .2 .1 .2 USAF Method 6 . 2 . 1 . 3 Torenbeek Method

P a r t V Contents Page iii

6 .2 .2 Commercial Transport Airplanes - 6 . 2 . 2 . 1 GD Method

6 . 2 . 2 . 2 Torenbeek Method ; 6.2 .3 M i l i t a r y P a t r o l , Bomb and Transport

Airplanes 6.2 .4 Fighter and Attack Airplanes

6 . 2 . 4 . 1 GD Method 6.3 PROPELLER WEIGUT ESTIMATION

6 . 3 . 1 General Aviation Airplanes 6 .3 .2 Commercial Transport Airplanes

6 .3 .2 .1 GD Method 6 .3 .2 .2 Torenbeek Method

6.3.3 M i l i t a r y P a t r o l , Bomb and Transport Airplanes

6 .3 .4 Figh te r and A t t a c k Airplanes 6.4 FUEL SYSTEM WEIGHT ESTIMATION

6 . 4 . 1 General Aviation Airplanes 6 . 4 . 1 . 1 Cessna Method 6 .4 .1 .2 USAF Method 6 .4 .1 .3 Torenbeek Method

6 .4 .2 Commercial Transport Airplanes 6 . 4 . 2 . 1 GD Method 6 .4 .2 .2 Torenbeek Method

6 . 4 . 3 M i l i t a r y P a t r o l , Bomb and Transport Airplanes

6.4 .4 Fighter and Attack Airplanes 6.5 PROPULSION SYSTEM WEIGHT ESTIMATION

6 . 5 . 1 General Aviation Airplanes 6.5. 1.1 Cessna Method 6 . 5 . 1 . 2 USAF Method 6.5 .1 .3 Torenbeek Method

6 .5 .2 Commercial Transport Airplanes 6 . 5 . 2 . 1 GD Method 6 .5 .2 .2 Torenbeek Method

6.5. 3 M i l i t a r y P a t r o l , Bomb and Transport Airplanes

6.5 .4 Figh te r and A t t a c k Airplanes

CLASS I1 METHOD FOR ESTIMATING FIXED EQUIPMENT WEIGHT 7 . 1 FLIGHT CONTROL SYSTEM WEIGHT ESTIMATION

7 . 1 . 1 General Aviation Airplanes 7 . 1 . 1 . 1 Cessna Method 7 . 1 . 1 . 2 USAF Method 7 .1 .1 .3 Torenbeek Method

7 .1 .2 Commercial Transport Airplanes 7 . 1 . 2 . 1 GD Method 7 . 1 . 2 . 2 Torenbeek Method

7 .1 .3 M i l i t a r y P a t r o l , Bomb and Transport Airplanes 7 . 1 . 3 . 1 GD Method

P a r t V Contents Page i v

P a r t V

7.1.4 Figh te r and A t t a c k Airplanes 100 7.1.4.1 GD Method 100

HYDRAULIC AND/OR PNEUMATIC SYSTEM WEIGHT ESTIMATION 101 ELECTRICAL SYSTEM WEIGHT ESTIMATION 101 7.3.1 General Aviation Airplanes 101

7.3.1.1 Cessna Method 101 7.3.1.2 USAF Method 101 7.3.1.3 Torenbeek Method 102

7.3.2 Commercial Transport Airplanes 102 7.3.2.1 GD Method 102 7.3.2.2 Torenbeek Method 102

7.3.3 M i l i t a r y P a t r o l , Bomb and Transport Airplanes 102 7.3.3.1 GD Method 102


WEIGHT ESTIMATION FOR INSTRUMENTATION, AVIONICS AND ELECTRON1 CS 103 7.4.1 General Aviation Airplanes 103

7.4.1.1 Torenbeek Method 103 7.4.2 Commercial Transport Airplanes 103

7.4.2.1 GD Method (Modified) 103 7.4.2.2 Torenbeek Method 104

7.4.3 M i l i t a r y P a t r o l , Bomb and Transport Airplanes 104

7.4.4 Figh te r and A t t a c k Airplanes 104 WEIGHT ESTIMATION FOR AIR-CONDITIONING, PRESSURIZATION, ANTI- AND DE- ICING SYSTEMS 104 7.5.1 General Aviation A i rp lanes 104

7.5.1.1 USAF Method 104 7.5.1.2 Torenbeek Method 104

7.5.2 Commercial Transport Airplanes 105 7.5.2.1 GD Method 105 7.5.2.2 Torenbeek Method 105

7.5.3 M i l i t a r y P a t r o l , Bomb and Transport Airplanes 105 7.5.3.1 GD Method 105

7.5.4 Figh te r and Attack Airplanes 105 7.5.4.1 GD Method 105

WEIGHT ESTIMATION FOR THE OXYGEN SYSTEM 106 7.6.1 General Aviation Airplanes 106 7.6.2 Commercial Transport Airplanes 106

7.6.2.1 GD Method 106 7.6.2.2 Torenbeek Method 106

7.6.3 M i l i t a r y P a t r o l , Bomb and Transport Airplanes -C - 106


AUXILIARY POWER UNIT WEIGH!l! ESTIMATION 107 FURNISHINGS WEIGHT ESTIMATION 107

Contents Page v

7.8.1 G e n e r a l A v i a t i o n A i r p l a n e s 7.8.1.1 C e s s n a Method - 7.8.1.2 T o r e n b e e k Method

- 7.8.2 C o m m e r c i a l T r a n s p o r t A i r p l a n e s

7. 8.2.1 GD Method 7.8.2.2 T o r e n b e e k Method

7.8.3 M i l i t a r y P a t r o l , Bomb and T r a n s p o r t A i r p l a n e s 7.8.3.1 GD Method

7.8.4 F i g h t e r and A t t a c k A i r p l a n e s 7.9 WEIGIiT ESTIMATION OF BAGGAGE AND CARGO

HANDLING EQUIPMENT 7.10 WEIGHT ESTIMATION OF OPERATIONAL ITEMS 7.11 ARMAMENT WEIGHT ESTIMATION 7.12 WEIGHT ESTIMATION FOR GUNS, LAUNCHERS

AND WEAPONS PROVISIONS 7.13 WEIGHT ESTIMATION OF FLIGHT TEST

INSTRUMENTATION 7.14 WEIGHT ESTIMATION FOR AUXILIARY GEAR 7.15 BALLAST WEIGHT ESTIMATION 7.16 ESTIMATING WEIGKT OF PAINT 7.17 ESTMATING WEIGET OF Wet,

8. LOCATING COMPONENT CENTERS OF GRAVITY 8.1 C.G. LOCATIONS OF STRUCTURAL COMPONENTS 8.2 C. G. LOCATIONS OF POWERPLANT COMPONENTS 8.3 C.G. LOCATIONS OF FIXED EQUIPMENT

9. CLASS I1 WEIGHT AND BALANCE ANALYSIS 117 9.1 EFFECT OF MOVING COMPONENTS ON OVERALL

AIRPLANE CENTER OF GRAVITY 117 9.2 EFFECT OF MOVING THE WING ON OVERALL AIRPLANE

CENTER OF GRAVITY AND ON OVERALL AIRPLANE AERODYNAMIC CENTER 119

10. CLASS I1 METHOD FOR ESTIMATING AIRPLANE INERTIAS 121

11. REFERENCES 123

APPENDIX A: DATA SOURCE FOR AIRPLANE COMPONENT WEIGHTS AND FOR WEIGHT FRACTIONS 125

APPENDIX B: DATA SOURCE FOR NONDIMENSIONAL R A D I I OF GYRATION FOR AIRPLANES 197

12. INDEX 207

Part v C o n t e n t s P a g e v i

TABLE OF SYMBOLS P P P I P I l l l P l P I P P P

FAR

h

i n t

Part v

wing aspect r a t i o ----- Hor. t a i l , Vert. t a i l o r ----- Canard aspect r a t i o I n l e t cap tu re a r e a per i n l e t f t 2

cons tant i n Eqn.(5.42) and Table 5.1

wing span f t Hor. t a i l , V e r t . t a i l o r Canard span f t cons tant i n Eqn. (5 .42) and Table 5.1

wing mean geometric chord f t

mean geometric chord of f t hor. t a i l , v e r t . t a i l o r canard cons tant i n Eqn. ( 5 . 4 2 ) and Table 5 .1

Drag c o e f f i c i e n t ----- L i f t c o e f f i c i e n t ----- Airplane l i f t - c u r v e s l o p e rad-'

Normal f o r c e c o e f f i c i e n t ----- cons tant i n Eqn.(5.42) and Table 5.1

P rope l l e r diameter f t

Used i n i n e r t i a ca lcs . f t

Federal A i r Regulation ----- a c c e l e r a t i o n of g r a v i t y f t l s e c 2

F l i g h t design g r o s s wht l b s

a l t i t u d e f t maximum fuse lage he ight f t *

-

f r a c t i o n of f u e l t anks which a r e i n t e g r a l ----- Moment of i n e r t i a s l u g s / f t 3

Symbols Page v i i

K - eons tan t as d e f i n e d i n equat ions below:

Kc ( 4 .6 ) K ( 6 .9 ) a n d (6 .10) b u t Kec (6 .23) n a t e t h a t values d i f f e r

Kf (5 .27 ) K f c f ( 7 . 9 ) K ( 5 .42 ) Kh ( 5 .19 ) g r

K i n l (5 .26) a n d (5 .28) Klav (7 .44 ) Km (6 .9 )

Kn (5 .29) Kosc (6 .38) K P

(6 .2 ) K W

(6 .4 )

'prop1 o r 2 (6 .13) or (6 .14 ) Kr (6 .11)

Ks (6 .12) Kth r ( 6 . 6 ) K, ( 5 . 2 0 ) Kw ( 5 . 9 ) . (5 .10)

Kst (7 .46)

K f SP

specific w e i g h t of f u e l l b s l g a l

K 4

G u s t a l l ev ia t ion fac tor , see Eqn. (4 .16)

f l e n g t h of f u s e l a g e f t

f-n l e n g t h of f u s e l a g e minus f t nacelle

Distance from wing l / 4 c

to 114ch,v,c f t

i n l e t l e n g t h from l i p t o f t c o m p r e s s o r face l e n g t h of passenger c a b i n f t

1 s h o c k s t r u t l e n g t h f o r main 'rn o r n g e a r or for nose g e a r f t

O v e r a l l a irplane l e n g t h f t i n l e t d u c t l e n g t h f t

Lr ramp l e n g t h f t

M Mach number

Mission f u e l f r a c t i o n none (Mff= End w e i g h t l B e g i n w e i g h t )

Load f a c t o r ----- Number of ( s e e s u b s c r i p t ) -----

~ a r t v Symbols Page v i i i

maximum fuse l . per imeter f t

design u l t . cabin press . p s i

required take-off power hp

maximum s t a t i c p ressu re a t engine compressor f a c e p s i

dynamic p ressu re psf

Range nm o r m Radius of i n e r t i a about f t x ,y ,z a x i s r e spec t ive ly

Non-dimensional r a d i u s ----- of i n e r t i a about x ,y ,z ax iz resp.

Wing a r e a f t 2

To ta l c o n t r o l s u r f a c e f t a r e a f r e i g h t f l o o r a r e a f t

Fuselage g r o s s s h e l l a r e a f t 2

or. . V e r t . o r Can. a r e a it2

Rudder a r e a f t 2

Shaf t horsepower h~

th ickness r a t i o ----- maximum r o o t th ickness f t

Derived g u s t v e l o c i t y f p s

True a i rspeed mph, f p s , k t s

Design maneuvering speed KEAS

Design speed f o r maxinum KEAS gus t i n t e n s i t y

Design c r u i s e speed KEAS -

Design d ive speed KEAS

Maximum l e v e l speed a t KEAS a t s e a l e v e l

Symbols Page i x P a r t V

Greek Symbols ' I I P I I I = = a P P P

a

Subscr ip t s ======I===

a i a p i

aps i

apu arm aux b a l bc b l C CC

cg c r crew C

P a r t V

Volume of passenger cabin f t 3

Vo1. of pass. and cargo f t 3

+lg s t a l l speed KE AS

maximum fuse lage width f t

Weight l b s Weight of component i l b s

d i s t a n c e from some ref. f t d i s t a n c e from some re f . of component i f t Distance from v e r t . t a i l root t o where h o t . is mounted on t h e v. t. f t

angle of a t t a c k of a i r p l a n e rad. downwash angle a t h. t. rad. a i r d e n s i t y s l u g s / i t 3 wing t a p e r r a t i o ----- t a p e r r a t i o f o r hor. t a i l , ve r t . t a i l o r canard --we-

a i r p l a n e mass r a t i o , s e e Eqn. (4 .17 )

sweep angle a t nth chord s t a t i o n

a i r induct ion a i r cond i t ion ing , p r e s s u r i z a t i o n , de- icing and an t i- ic ing system accessory d r i v e s , powerplant c o n t r o l s , s t a r t i n g and i g n i t i o n system a u x i l i a r y power u n i t armament a u x i l i a r y b a l l a s t baggage and cargo handling equipment b lades canard cabin crew cen te r of g r a v i t y crew crew Cruise

Symbols Page x

D e ec e l s emp eng e s s e t c

E f f c f d f dc f e q f l . boat f s f t i f u r F 9 glw h bps H i i a e

in f l r e f i n 1 l i m L L LE m max MZF n neg OPS OSC OX

Pax P P PC POS Prop P t Pwr PL

P a r t V

Dive engines ( a l l 1 1 engine c o n t r o l s e l e c t r i c a l system empennage engine (one only) engine s t a r t i n g system e t c e t e r a (p lease pronounce a s eTcetera

and as ekcetera) Empty fuse lage f l i g h t con t ro l system f u e l dumping system f l i g h t deck crew f ixed equipment f l y i n g boat f u e l system f l i g h t t es t i n s t r u m n t a t i o n fu rn i sh ings Mission f u e l landing gear guns, launchers and weapons provis ions hor izon ta l t a i l hydraul ic and pneumatic system maximum l e v e l f l i g h t a t s e a l e v e l instrumentat ion instrumentat ion, av ion ics and e l e c t r o n i c s i n- f l i g h t r e f u e l l i n g system i n l e t (s) l i m i t Landing ( s u b s c r i p t t o W ) maximum dive ( s u b s c r i p t t o V) Leading Edge maximum maximum Maximum zero f u e l n a c e l l e negat ive opera t iona l items o i l system and o i l cooler oxygen system passengers p r o p e l l e r s ( s u b s c r i p t t o N) propulsion system ( s u b s c r i p t t o W) p rope l l e r c o n t r o l s p o s i t i v e 2 -

p r o p e l l e r p a i n t powerplant Payload

Symbols Page x i

ramp sprckr s t r u c t SUPP - t - t f o t r t r o o p TO u l t u l t . 1 . v W wb w i X X D Y Y D Zz

Acronyms ========

APU C.G., C*g* OWE TBP

r a w supercharger s t r u c t u r e bladder support s t r u c t u r e f u e l tanks trapped f u e l and o i l t h r u s t r eve r se r system t roop(s1 Take-of f u l t ima te u l t i m a t e landing v e r t i c a l t a i l wing wing + body water i n j e c t i o n system about x-, y-, z- axis r e spec t ive ly

Auxil iary power u n i t Center of g r a v i t y Operating weight empty Turboprop

P a r t V Symbols Page x i i

ACKNOWLEDGEMENT PPP=PPPPPPPIP=I

Writing a book on a i r p l a n e weight es t imat ion is impossible without t h e supply of a l a r g e amount of data . The author is g r a t e f u l t o t h e following companies f o r supplying t h e weight da ta , t h e weight manuals and t h e weight est imating procedures which made t h e book what it is :

Beech A i r c r a f t Corporation Boeing Commercial Airplane Company Canadair Cessna A i r c r a f t Company DeHavilland A i r c r a f t Company of Canada Gates L e a r j e t Corporation Gulfstream Aerospace Corporation Lockheed A i r c r a f t Corporation McDonnell Douglas Corporation NASA, Ames Research Center Rinaldo Piaggio S.p.A. Rockwell I n t e r n a t i o n a l Royal Netherlands A i r c r a f t Factory, Fokker SIAI Marchett i S.p.A.

A s i g n i f i c a n t amount of a i r p l a n e design information has been accumulated by t h e author over many yea r s from t h e following magazines:

I n t e r a v i a (Swiss, monthly) F l i g h t I n t e r n a t i o n a l ( B r i t i s h , weekly) Business and Commercial Aviation (USA, monthly) Aviation Week and Space Technology (USA, weekly Journal of A i r c r a f t (USA, AIAA, monthly)

The author wishes t o acknowledge t h e inpor tan t r o l e played by t h e s e magazines i n h i s own development a s an ae ronau t i ca l engineer. Aeronautical engineering s tuden t s and graduates should read t h e s e magazines regular ly .

P a r t V Page x i i i

Part V ':u Page x i v

1. INTRODUCTION P P I P P P O P I I P P P P P

The purpose of t h i s series of books on Airplane Design is t o f a m i l i a r i z e aerospace engineering s tuden t s with t h e design methodology and design dec i s ion making involved i n t h e process of designing a i rp lanes .

The series of books is organized as follows:

PART I: PRELIMINARY SIZING OF AIRPLANES PART 11: PRELIMINARY CONFIGURATION DESIGN AND

INTEGRATION OF THE PROPULSION SYSTEM PART 111: LAYOUT DESIGN OF COCKPIT, FUSELAGE, WING

AND EMPENNAGE: CUTAWAYS AND INBOARD PROFILES

PART I V : LAYOUT DESIGN OF LANDING GEAR AND SYSTEMS PART V: COMPONENT WEIGRI' ESTIMATION PART V I : PRELIMINARY CALCULATION OF AERODYNAMIC,

THRUST AND POWER CHARACTERISTICS PART V I I : DETERMINATION OF STABILITY, CONTROL AND

PERFORMANCE CHARACTER1 STICS : FAR AND MILITARY REQUIREMENTS

PART V I I I : AIRPLANE COST ESTIMATION: DESIGN, DEVELOPMENT, MANUFACTURING AND OPERATING

The purpose of PART V is t o p resen t methods f o r est imating a i r p l a n e component weights and a i r p l a n e iner t ias during a i r p l a n e preliminary design.

Two methods are presented: they are c a l l e d t h e Class I and t h e Class I1 method respect ive ly .

The Class I method relies on t h e es t imat ion of a percentage of t h e f l i g h t design g ross weight (= take-off weight f o r most a i r p l a n e s ) of major a i r p l a n e components. These percentages are obtained from actual weight d a t a for e x i s t i n g a i rp lanes . The usual procedure is t o average t h e s e percentages f o r a number of a i r p l a n e s similar t o t h e one being designed. These averaged percentages are mul t ip l i ed by t h e take-off weight t o obta in a f i r s t estimate of t h e weight of each major component.

The method can be used with minimal knowledge about t h e a i r p l a n e being designed and requ i res very l i t t l e engineer ing work. However, t h e accuracy of t h i s method is l imi ted . It should be used only i n a s s o c i a t i o n with pre- l iminary design sequence I as ou t l ined i n P a r t 11 @ee S tep l o , p.15). -

Chapter 2 p r e s e n t s t h e Class I method f o r es t imat ing

P a r t V Chapter 1 Page 1

a i r p l a n e component weights i n t h e form of a step- by- step precedure. Three example a p p l i c a t i o n s a r e a l s o given.

- Chapter 3 p r e s e n t s a C l a s s I method f o r e s t ima t ing

a i r p l g n e moments of i n e r t i a . Example a p p l i c a t i o n s a r e a l s o given.

C l a s s I1 methods a r e based on weight equa t ions f o r more d e t a i l e d a i r p l a n e components and groupings. These equa t ions have a s t a t i s t i c a l b a s i s . They do a l low t h e des igne r t o account f o r f a i r l y d e t a i l e d c o n f i g u r a t i o n des ign parameters. To use t h i s method it is necessary t o have a V-n diagram, a p re l imina ry s t r u c t u r a l arrangement and t o have decided on a l l systems which a r e needed f o r t h e ope ra t ion of t h e a i r p l a n e under s tudy.

The C l a s s I1 method should be used i n con junc t ion wi th p re l imina ry des ign sequence I1 as o u t l i n e d i n P a r t I1 (See S t ep 2 1 , p.19).

Chapter 4 p r e s e n t s t h e Class 11 method f o r e s t ima t ing a i r p l a n e component weights i n t h e form of a s tep- by- step procedure. A method f o r c o n s t r u c t i o n of a V-n diagram is included. Example a p p l i c a t i o n s a r e given.

A s p a r t of t h e C l a s s I1 weight e s t i m a t i o n procedure t h e a i r p l a n e empty weight is s p l i t i n t o t h r e e major groupings:

1. S t r u c t u r e weight 2. Powerplant weight 3. Fixed equipment weight

Chapters 5, 6 and 7 p r e s e n t t h e d e t a i l e d methodologies used i n determining t h e component weights w i th in each of t h e s e t h r e e groupings.

Chapter 8 c o n t a i n s d a t a and methods f o r r a p i d l y determining t h e c.g. l o c a t i o n of i n d i v i d u a l components.

A Class I1 method f o r performing a weight and ba lance a n a l y s i s is d i scussed i n Chapter 9.

Chapter 1 0 p r e s e n t s a Class I1 method f o r computing a i r p l a n e moments and produc ts of i n e r t i a .

Appendix A c o n t a i n s a d a t a base f o r a i r p l a n e component weights and weight f r a c t i o n s .

Appendix B c o n t a i n s a d a t a base f o r non-d imns iona l r a d i i of g y r a t i o n f o r a i r p l a n e s .


2. CLASS I METHOD FOR ESTIMATING AIRPLANE COMPONENT ----- - - - - - P = = = P = P I 3 = I I P = = = = = = P t = P P P I I = t P = = 3 L P P = ~ = w = = P = x = = m

WEIGHTS =======

The purpose of t h i s chap te r is t o prov ide a methodology f o r r a p i d l y e s t ima t ing a i r p l a n e component weights. The emphasis is o c r ap id and on spending as few engineer ing manhours as poss ib l e . Methods which f i t meet t h e s e o b j e c t i v e s a r e r e f e r r e d t o a s C l a s s I methods. They a r e used i n conjunc t ion wi th t h e f i r s t s t a g e i n t h e p re l imina ry des ign process , t h e one r e f e r r e d t o as 'p.d. sequence I ' i n P a r t I1 (See S t e p lo, p.15).

The Class I weight e s t ima t ing method r e l i e s on t h e assumption, t h a t w i th in each a i r p l a n e ca tegory it is p o s s i b l e t o exp res s t h e weight of major a i r p l a n e components ( o r groups) a s a s imple f r a c t i o n of one of t h e fo l lowing weights:

1. Gross take- off weight, WTO

2. F l i g h t des ign g r o s s weight, GW

3. Empty weight, WE

The reader is a l r eady f a m i l i a r wi th t h e d e f i n i t i o n of WTO and WE. The f l i g h t des ign g r o s s weight, GW is

t h a t weight a t which t h e a i r p l a n e can s u s t a i n i t s design u l t i m a t e load f a c t o r , nult. For c i v i l a i r p l a n e s GW and

W ~ o a r e o f t e n t h e same, a l though t h e r e a r e except ions .

For m i l i t a r y a i r p l a n e s GW and WTO a r e f r e q u e n t l y q u i t e d i f f e r e n t .

I n t h i s book, a l l component weight f r a c t i o n s are g iven r e l a t i v e t o t h e f l i g h t des ign g r o s s weight, GW. I n t h e component weight and weight f r a c t i o n d a t a presen- t e d i n Appendix A, both GW and WTO a r e l i s t e d f o r a l l

a i r p l a n e s f o r which d a t a a r e p resen ted .

S ince WTO is known from t h e p re l imina ry s i z i n g work

desc r ibed i n P a r t I, t h e va lue of GW can be e s t a b l i s h e d . -

The weight of any major a i r p l a n e component o r group can now be found r a p i d l y through m u l t i p l i c a t i o n of GW by


an appropr ia t e weight f r a c t i o n . For t h i s reason, t h e Class I weight method is a l s o r e f e r r e d t o a s t h e 'weight f r a c t j o n ' method.

gec t ion 2.1 p resen t s a step-by-step procedure f o r using weight f r a c t i o n s t o es t imate t h e component weight breakdown of a i rp lanes .

Sec t ion 2.2 p r e s e n t s example a p p l i c a t i o n s t o t h r e e a i rp lanes .

2 .1 A METIIQD FOR ESTIMATING C- - I n t h i s s e c t i o n t h e Class I method f o r est imating

a i r p l a n e component weights is presented i n t h e form of a step-by-step procedure.

SteD L i s t t h e following o v e r a l l weight va lues f o r t h e a i rp lane :

1. Gross take-off weight, WTO

2. Empty weight, WE

3. Mission Fuel Weight, WF

4. Payload weight. WpL

5. Crew weight, Wcr,

6. Trapped f u e l and o i l weight, W t f o

7. F l i g h t design g r o s s weight, GW

Weight items 1-6 a r e a l ready known from t h e prel iminary s i z i n g process descr ibed i n P a r t I (See Chapter 2 ) .

For most a i r p l a n e s , WTO and GW are t h e same. I n

t h e case of many m i l i t a r y a i r p l a n e s t h e r e is a d i f fe rence. Appendix A conta ins t a b l e s with a i r p l a n e weight d a t a on b a s i s of which a dec i s ion can be made about t h e r a t i o between WTO and GW. Sometimes t h e mission speci-

f i c a t i o n w i l l include t h i s information.

SteD 2: Proceed t o Appendix A and determine which a i r p l a n e category b e s t f i t s t h e a i r p l a n e which is being designed. I d e n t i f y those


a i r p l a n e s which w i l l be used i n est imating t h e weight f r a c t i o n s f o r t h e a i r p l a n e which is being designed.

S f e D 3: Make a l ist of t h e s i g n i f i c a n t a i r p l a n e components f o r which weights need t o be estimated. This l i s t w i l l vary some from one a i r - plane type t o t h e o the r . In many cases cer- t a i n weight items a r e a l ready s p e c i f i e d i n t h e mission s p e c i f i c a t i o n .

A t y p i c a l Class I component weight list con ta ins t h e following items:

1. Wing 2. Empennage

2 .1 Horizontal t a i l and/or canard 2.2 V e r t i c a l t a i l and/or canard

3. Fuselage (and/or tailbooms) 4. Nacelles 5. Landing gear

5.1 Nose gear 5.2 Main gear 5.3 T a i l gear 5.4 Outrigger gear 5.5 F loa t s

1. Engine(s1, t h i s may include a f t e rburne r s o r t h r u s t r eve r se r s

2. A i r induct ion system 3. P rope l l e r ( s1 4. Fuel system 5. Propulsion system

Fixed vWeiaht, W f e q

P a r t V

F l i g h t con t ro l system Hydraulic and pneumatic system E l e c t r i c a l system Instrumentat ion, av ion ics and e l e c t r o n i c s Air condi t ioning, p ressu r i za t ion , an t i- ic ing and de- icing system Oxygen system Jt- -

Auxil iary power u n i t Furnishings Baggage and cargo handling equipment Operat ional items

Chapter 2 Page 5

11. Armament 1 2 . Guns, l aunchers and weapons p r o v i s i o n s

- 13. F l i g h t tes t in s t rumen ta t ion 1 4 . Aux i l i a ry g e a r - - 1 5 . B a l l a s t 16 . P a i n t 1 7 . Other weight items n o t l i s t e d above

Consul t t h e miss ion s p e c i f i c a t i o n a s w e l l a s t h e a p p r o p r i a t e t a b l e s i n Appendix A f o r any weight i tems n o t l i s ted above.

The a i r p l a n e empty weight, WE is expressed as:

Whether o r n o t it is necessary t o s p l i t weight groupings I1 and 111 i n as many components as l i s t e d above depends on t h e expected e f f e c t of t h e s e components on t h e accuracy of t h e a i r p l a n e c.g. l o c a t i o n .

U s e as much d e t a i l a s necessary f o r r e a l i s m i n t h e C la s s I weight and balance a n a l y s i s of Chapter 1 0 , P a r t 11.

fie^ 4: From t h e a p p r o p r i a t e T a b l e ( s ) i n Appendix A dec ide on t h e weight f r a c t i o n s t o be used.

Frequent ly it w i l l be s u f f i c i e n t t o u se average f r a c t i o n va lues ob ta ined from a number of a i r p l a n e s wi th miss ions n o t t o o much d i f f e r e n t from t h e miss ion of t h e a i r p l a n e being designed. The r eade r should f a m i l i a r i z e himself wi th what t h e a i r p l a n e s f o r which weight f r a c t i o n d a t a a r e a v a i l a b l e , look l i k e and what t h e i r miss ions were. Th i s can be done by r e f e r r i n g t o J a n e ' s A l l t h e World A i r c r a f t (Ref. 81.. J a n e ' s c o n t a i n s an index i d e n t i f y i n g which i s s u e of J a n e ' s c o n t a i n s d e s c r i p t i o n s of c e r t a i n t y p e s of a i r p l a n e s .

I t is of g r e a t importance t o observe whether o r no t :

1. an a i r p l a n e has a s t r u t t e d (b raced ) wing 2. an a i r p l a n e is p r e s s u r i z e d 3. t h e land ing g e a r is mounted on t h e f u s e l a g e o r on

t h e wing 4. t h e engines a r e mounted on t h e wing o r f u s e l a g e

The reader should no te , t h a t most weight and weight f r a c t i o n d a t a i n Appendix A a r e f o r a i r p l a n e s wi th l a r g e l y aluminum primary s t r u c t u r e s . I f t h e a i r p l a n e being designed w i l l have t o c o n t a i n a s i g n i f i c a n t amount


of primary s t r u c t u r e made from composites, from lithium-aluminum o r from o t h e r m a t e r i a l s , it w i l l be necessary t o modify t h e weight f r a c t i o n s . Table 2.16, p.48, P a r t I may be u s e f u l i n t h i s regard.

A f t e r t h u s 'massaging' t h e weight f r a c t i o n d a t a , l ist t h e weight f r a c t i o n s t o be used. Make c a r e f u l n o t e s of reasons why s p e c i f i c f r a c t i o n s were s e l e c t e d .

e D 5 : Mul t ip ly t h e s e l e c t e d weight f r a c t i o n s by t h e GW va lue of S t e p 1 and l is t a l l s i g n i f i - c a n t a i r p l a n e component weights.

The Class I component weight d a t a t h u s ob ta ined a r e used i n t h e C la s s I weight and ba lance a n a l y s i s desc r ibed i n Chapter 10 of P a r t 11.

To i l l u s t r a t e t h e use of t h i s procedure , t h r e e examples a r e p resen ted i n Sec t ion 2.2.

e~ 6: Document t h e d e c i s i o n s made under S t e p s 1 through 5 i n a b r i e f , d e s c r i p t i v e r e p o r t .

I n t h i s s e c t i o n , t h r e e example a p p l i c a t i o n s of t h e Class I component weight e s t ima t ing method w i l l be d i scussed :

2 . 2 . 1 Twin Engine P r o p e l l e r Driven Airplane: Se l ene 2.2.2 Jet Transpor t : Ourania 2.2.3 F i g h t e r : E r i s

S f e D 1: O v e r a l l weight va lues f o r t h i s a i r p l a n e were determined as a r e s u l t of t h e p re l imina ry s i z i n g performed i n P a r t I. These weight v a l u e s a r e summarized i n sub- sub- section 3.7.2.6, P a r t I , p.178:

w ~ o 7,900 l b s WE = 4,900 l b s

WF = 1,706 l b s w~~ = 1 , 2 5 0 l b s ( P a r t I , p.49)

W t f o = 4 4 l b s makes up t h e balance.

The crew weight is included i n t h e payload o k t h i s a i r p l a n e . I t w i l l be assumed t h a t GW - WTO. Th is is

c o n s i s t e n t wi th t h e d a t a i n Tables A3.1 and A3.2.


F o r easy reference t h e airplane w i l l b e r e f e r r e d t o as the Selene, t h e name o f t h e Greek Moon Goddess .

- S t e ~ 2: T a b l e s A3.1 and A3.2 contain component

w e i g h t d a t a for airplanes i n t h e same category as t h e S e l e n e . S p e c i f i c a l l y , t h e fol lowing airplanes h a v e c o m p a r a b l e s izes a n d missions: C e s s n a 310C, Beech 65 Queen A i r , C e s s n a 404-3 and C e s s n a 414A.

S t e ~ 3: F o r reasons of b r e v i t y , only t h e f o l l o w i n g component w e i g h t s are c o n s i d e r e d :

Wing Empennage F u s e l a g e Nacelles Land ing Gear Power P l a n t F i x e d Eqpmt

B e D 4 : T h e f o l l o w i n g t a b l e lists t h e p e r t i n e n t w e i g h t f ract ions a n d t h e i r a v e r a g e d values. B e c a u s e t h e i n t e n t is t o apply conventional metal cons t ruc t ion methods t o t h e S e l e n e t h e r e is no reason t o a l t e r t h e a v e r a g e d w e i g h t fractions.

Beech C e s s n a C e s s n a C e s s n a S e l e n e 65 QA 310C 404-3 414A A v e r a g e

Pwr PltIGW 0 ,219 0.259 0 .194 0 ,206 0 .220 F i x EqpIGW 0.123 0 .103 0 .134 0.167 0 .132 Empty W h t I G W 0 .638 0 .628 0 .596 0.665 0 . 6 3 1

Wing G r p I G W 0.091 0.094 0 .102 0.094 0 .095 Emp, G r p I G W 0.021 0.024 0.022 0 ,024 0 .023 Fus . G r p I G W 0.082 0 .066 0 .073 0 ,100 0 .080 Nac. G r p I G W 0.039 0.027 0 .034 0.029 0.032 Gear G r p l G ~ 0.060 0.054 0 .038 0 .045 0 .049

Note t h a t t h e r a t i o of WE/GW wh ich f o l l o w s from t h e

preliminary s i z i n g , is 4 ,90017 ,900 = 0.62. T h i s is close t o t h e average value of 0 .631 i n t h e a b o v e t a b u l a t i o n .

S f e D 5: Using t h e a v e r a g e d weight f r ac t ions from S t e p 4 , t h e f o l l o w i n g prel iminary component w e i g h t summary can b e d e t e r m i n e d :

P a r t v C h a p t e r 2 P a g e 8

Selene Component F i r s t weight Adjustment Class I Class I

es t imate weight weight (a lum. ) (compos.)

l b s l b s l b s l b s ........................................................ Wing 751 -13 7 3 8 627 Empennage 182 - 3 1 7 9 1 5 2 Fuselage 632 -11 62 1 5 2 8 Nacelles 253 - 4 2 4 9 212 Landing Gear 3 87 - 7 3 80 3 80 Power P lan t 1 , 7 3 8 -30 1 , 7 0 8 1 , 7 0 8 Fixed Eqp. 1 , 0 4 3 -1 8 1 , 0 2 5 1 , 0 2 5

Empty Wht 4 , 9 8 6 - 86 4 ,900 4 , 6 3 2 ........................................................ Payload 1 , 2 5 0 1 , 2 5 0 Fuel 1 , 7 0 6 1 , 7 0 6 Trapped f u e l and o i l 44 4 4 ........................................................ Take-off Gross Weight 7 ,900 7 , 6 3 2

When t h e numbers i n t h e f i r s t column a r e added, they y i e l d an empty weight of 4 , 9 8 6 l b s ins t ead of t h e des i red 4 , 9 0 0 lbs . The d i f f e r e n c e is due t o round-off e r r o r s i n t h e weight f r a c t i o n s used. It is b e s t t o ' d i s t r i b u t e ' t h i s d i f f e r e n c e over a l l items i n propor t ion t o t h e i r component weight value l i s t e d i n t h e f i r s t column.

For example, t h e wing adjustment number is a r r i v e d a t by mult iplying 86 l b s by 751/4986.

It is q u i t e p o s s i b l e t h a t i n o t h e r a i r p l a n e s t h e adjustment w i l l t u r n out t o be p o s i t i v e ins tead of negat ive.

I f t h e judgement is made t o manufacture t h e Selene with composites a s t h e primary s t r u c t u r a l ma te r i a l s s i g- n i f i c a n t weight savings can be obtained. A conservat ive assumption is t o apply a 1 5 percent weight reduct ion t o wing, empennage, fuse lage and nace l l e s . The r e s u l t i n g weights a r e a l s o shown i n t h e Class I weight t abu la t ion . Note t h e reduct ion i n empty weight of 268 lbs . Using t h e weight s e n s i t i v i t y aWTO/aWE = 1 .66 as computed i n

sub-sub-section 2.7.3.1 i n P a r t I, an o v e r a l l reduct ion i n WTO of 1 . 6 6 ~ 2 6 8 - 545 i b s can be achieved. -C -

The designer has t h e obvious choice t o f l y t h e same mission with (545 - 268) = 277 l b s l e s s f u e l o r t o simply add t h e 545 l b s t o t h e use fu l load of t h e Selene.


The component weight v a l u e s i n t h e column l a b e l l e d : ' C l a s s I weight (alum. 1 ' a r e t h o s e t o be used i n t h e C l a s s - I weight and balance a n a l y s i s of t h e Selene. This corresponds t o S t e p 1 0 a s o u t l i n e d i n Chapter 2 , P a r t 11. The Class I weight and balance a n a l y s i s f o r t h e Se l ene is c a r r i e d o u t i n Chapter 10 of P a r t I1 (See pp. 246-250) .

;Ete~ 6: To save space , t h i s s t e p has been omit ted.

2.2.2 Je t

a e ~ 1: Overa l l weight va lues f o r t h i s a i r p l a n e were determined a s a r e s u l t of t h e p re l imina ry s i z i n g performed i n P a r t I. These weight v a l u e s are summarized i n sub- sub- section 3.7.3.6. P a r t I , p. 183:

w ~ o = 127,000 l b s WE = 6 8 , 4 5 0 l b s

WF = 25 ,850 l b s W~~ = 30 ,750 l b s ( P a r t I, p.54)

W t f o = 925 l b s 'crew = 1 , 0 2 5 l b s ( P a r t I , p.58)

I t w i l l be assumed t h a t GW = WTO f o r t h i s a i r p l a n e .

This is c o n s i s t e n t wi th t h e d a t a i n Tables A 7 . 1 through A 7 . 5 .

For easy r e f e rence t h e a i r p l a n e w i l l be r e f e r r e d t o a s t h e Ourania, t h e name of t h e Greek Muse of Astronomy.

=D 2: Tables A 7 . 1 through A 7 . 5 c o n t a i n component weight d a t a f o r a i r p l a n e s i n t h e same ca tegory as t h e Ourania. S p e c i f i c a l l y t h e fo l lowing a i r p l a n e s have comparable s i z e s and miss ions: McDonnell-Douglas DC-9-30 and MD-80, Boeing 737-200 and 727-100.

SfeD 3 : For reasons of b r e v i t y , o n l y t h e fol lowing component weights are considered:

Wing Empennage Fuselage Nace l les Landing Gear Power P l a n t Fixed Eqpmt

The fol lowing t a b l e lists t h e p e r t i n e n t weight f r a c t i o n s and t h e i r averaged va lues . Because t h e i n t e n t is t o apply convent iona l metal c o n s t r u c t i o n methods t o t h e Ourania, t h e r e is no reason t o a l t e r t h e averaged weight f r a c t i o n s .

P a r t V Chapter 2 Page 1 0

McDonnell-Douglas Boeing Our a n i a DC-9-30 MD- 80 737-200 727-100 Average

Pwr P l t IGW 0.076 0.079 0.071 0.078 0.076 Fix EqpIGW 0.175 0.182 0.129 0.133 0.155 Empty~ht IGW 0.538 0.564 0.521 0.552 0.544

Wing Grp/GW 0.106 0.111 0.092 0.111 0.105 Emp. Grp/GW 0.026 0.024 0.024 0.026 0.025 Fus. GrpIGW 0.103 0.115 0.105 0.111 0.109 Nac. Grp/GW 0.013 0.015 0.012 0.024 0.016 GearGrpIGW 0.039 0.038 0.038 0.045 0.040

Note t h a t t h e r a t i o of WE/GW which fo l lows from t h e

p re l imina ry s i z i n g , is 68,450/127,000 = 0.539. This is c l o s e t o t h e average va lue of 0.544 i n t h e above t a b u l a t i o n .

s e r , 5 : Using t h e averaged weight f r a c t i o n s j u s t determined, t h e fol lowing p re l imina ry component weight summary can be determined:

Ourania Component F i r s t weight Adjustment Class I Class I

e s t i m a t e weight weight (alum.) ( l i l a l u m . )

l b s l b s l b s l b s ........................................................ Wing 13,335 +3 2 9 13,664 12,298 Empennage 3,175 + 78 3,253 2,928 Fuselage 13,843 +341 14,184 12,766 Nacelles 2,032 + 50 2,082 1,874 Landing Gear 5,O 80 +I25 5,205 5,205 Power P l a n t 9,652 +2 39 9, 891 9, 891 Fixed Eqp. 19,685 +4 86 20,171 20,171 ........................................................ Empty Wht 66,802 +1,648 68,450 65,133 ........................................................ Payload 30,750 30,750 Crew 1,025 1,025 Fuel 25,850 25,850 Trapped f u e l and o i l 925 925 ........................................................ Take-off Gross Weight 127,000 123,683

When t h e numbers i n t h e f i r s t column a r e addeU, t hey y i e l d an empty weight of 66,802 l b s i n s t e a d of t h e d e s i r e d 68,450 l b s . The d i f f e r e n c e is due t o round-off e r r o r s i n t h e weight f r a c t i o n s used. It is b e s t t o ' d i s t r i b u t e ' t h i s d i f f e r e n c e over a l l i tems i n p ropor t ion


t o t h e i r component weight values l i s t e d i n t h e f i r s t column. -

P o r example, t h e wing adjustment number is a r r ived a t b f m u l t i p l y i n g 1 , 6 4 8 l b s by 13,335/66,802. When s o doing, t h e sum of t h e adjusted component weights is s t i l l 4 1 l b s shy of t h e des i red goal. T h a t new d i f f e r e n c e is then r e d i s t r i b u t e d i n t h e same manner.

I t w i l l be noted t h a t t h e adjustments h e r e a r e p o s i t i v e whereas f o r t h e l i g h t twin they were negative. I t a l l depends on t h e weight f r a c t i o n roundoffs, how t h i s comes out .

I f t h e judgement is made t o manufacture t h e Ourania with lithium/aluminum a s t h e primary s t r u c t u r a l ma te r i a l , s i g i f i c a n t weight savings can be obtained. A reasonable assumption is t o apply a 1 0 percent weight reduct ion t o wing, empennage, fuse lage and nace l les . The r e s u l t i n g weights a r e a l s o shown i n t h e Class I weight t abu la t ion . Note t h e reduct ion i n empty weight of 3 , 3 1 7 l b s . Using t h e weight s e n s i t i v i t y a W T O l a W E = 1 .93 a s computed i n

sub-sub-section 2.7.3.2 i n P a r t I, an o v e r a l l reduct ion i n WTO of 1.93x3.317 - 6 ,402 l b s can be achieved.

The designer has t h e obvious choice t o f l y t h e same mission with ( 6 , 4 0 2 - 3 , 3 1 7 ) = 3 ,085 l b s l e s s f u e l o r t o add t h e 6 ,402 l b s t o t h e use fu l load of t h e Ourania.

The component weight va lues i n t h e column labe l l ed : 'Class I weight (alum. ' a r e those t o be used i n t h e Class I weight and balance a n a l y s i s of t h e Ourania. T h i s corresponds t o S tep 10 a s ou t l ined i n Chapter 2 , P a r t 11. The Class I weight and balance a n a l y s i s of t h e Ourania is c a r r i e d out i n Chapter 1 0 of P a r t I1 (See pp. 250-254.

UD 6: To save space, t h i s s t e p is omitted.

Ster, Overa l l weight values f o r t h i s a i r p l a n e were determined a s a r e s u l t of t h e prel iminary s i z i n g performed i n P a r t I. These weight va lues a r e summarized i n sub-sub-section 3 .7 .4 .5 , P a r t I , p.191:

Wp = 18,500 l b s WpL = 12 ,000 l b s ( P a r t I, p.60)

W t f o = 300 l b s 'crew = 200 l b s ( P a r t I , p.66)


I t w i l l b e assumed t h a t GW - 0.95WT0 f o r t h i s a i r -

plane. T h i s is c o n s i s t e n t w i th t h e d a t a i n Tab l e s A9.1 th rough A9.6.

For ea sy reference t h e airplane w i l l b e r e f e r r e d t o as t h e E r i s , t h e name of t h e Greek Goddess of War.

When look ing u p t h e actual bomb weight f o r a nominal 500 l b s bomb, it w i l l b e d i s cove red t h a t t h i s weight is 531 l b s and no t 500 l b s . Tha t is a d i f f e r e n c e of 20x31 = 6201bs. On t h e o t h e r hand, t h e normal ammunition f o r t h e s t a n d a r d GAU-8A gun drum weighs 1 ,785 and no t 2,000 l b s . The d i f f e r e n c e is -215 l b s . The actual payload is t h e r e- f o r e 405 l b s more t h a n o r i g i n a l l y planned.

Tab l e s A9.1 th rough A9.6 contain component weight d a t a f o r airplanes i n t h e same ca t ego ry as t h e E r i s . S p e c i f i c a l l y t h e fo l lowing airplanes have comparable s izes and miss ions : Republ ic FIOSB, Vought F SU, and Grumman A2F.

WD 3: For reasons of b r e v i t y o n l y t h e fo l lowing component weigh ts are cons idered :

Wing Empennage Fuse lage Eng. Sec t . Landing Gear Power P l a n t Fixed Eqpmt

S f e ~ 4: The fo l lowing t a b l e lists t h e per t inen t weight f r a c t i o n s and t h e i r averaged values. S i n c e E r i s w i l l b e made from conventional a luminum materials, t h e r e is no reason t o a l te r t h e averaged weight f r a c t i o n s .

Republ ic Vought Grumman E r is F105B F 8U A2F (A6 1 Average

Pwr P l t /GW 0.246 0.257 0.162 0.222 P ix EqpIGW 0.155 0.135 0.159 0.150 Empty W h t / G W 0.797 0.722 0.651 0.723

Wing Grp /GW 0.109 Emp. G r p I G W 0.031 FUS. Grp /GW 0.187 Eng.Sect./Gw 0.003 Gear G r p I G W 0.059 Engine ( s 1 /GW 0.197

9.6 N.A Use: 1 2

GW/WTO 0.92 0.79 1.0 Use: 0.95


Note: a l l f r a c t i o n d a t a were b a s e d on GW w i t h o u t ex- t e r n d stores!

Wote t h a t t h e r a t i o o f WE/GW wh ich f o l l o w s from t h e

prel iminary s i z i n g , is 3 3 , 5 0 0 / 5 4 , 5 0 0 = 0.615. T h i s is lower t h a n t h e a v e r a g e value of 0 .723 i n t h e a b o v e tabula t ion . T h e reason is t h a t t h e d a t a b a s e is f o r o l d e r f i g h t e r s , two of which are USN f i g h t e r s . A l s o note t h e l a r g e va lue nult fo r t h e FlOSB. .

s f e ~ 5: Using t h e a v e r a g e d w e i g h t f r a c t i o n s j u s t d e t e r m i n e d , t h e f o l l o w i n g preliminary component w e i g h t summary can b e d e t e r m i n e d :

E r is Component F i r s t w e i g h t A d j u s t m e n t C l a s s I

estimate weight (alum.

l b s l b s l b s ................................................. Wing 6 ,922 - 160 6 ,762 Empennage 1 , 6 3 5 - 3 8 1 ,597 F u s e l a g e 7 , 5 2 1 -174 7 ,347 Eng . S e c t . 164 - 4 1 6 0 Land ing Gear 2 ,834 - 66 2 , 7 6 8

Power p l an t 12 ,099 predicted f r o m f rac t ion d a t a E n g i n e s 9 ,265 p r e d i c t e d f r o m f rac t ion d a t a E n g i n e s 6 ,000 actual f o r F404 .s w i t h A/B

E n g i n e s 6 , 0 0 0 Eng . S e c t . 12 ,099- 9,265 = 2 ,834

F i x . Eqpmt 8 ,175 predicted from f r a c t i o n d a t a Ammo 2 ,000 ( o r i g i n a l e s t i m . ) F i x . Eqpmt -Ammo 6 ,175 -143 = 6 , 0 3 2

GAU-8A Gun ( A c t u a l w e i g h t ) 2 ,014 F i x . Eqpmt-Gun 4 ,018 ................................................. Empty Wht 39 ,350 -5 85 33 ,500 ................................................. P i l o t 200 P a y l o a d : ammo 1 , 7 8 5

: bombs 1 0 , 6 2 0 T r a p p e d f u e l and o i l 300 F u e l 1 8 , 5 0 0 ................................................. Take- of f Gross Weight 64 ,905

When t h e numbers i n t h e f i r s t column are a d d e d , t h e y y i e l d an empty w e i g h t o f 39 ,350 l b s i n s t e a d o f t h e

P a r t V C h a p t e r 2 P a g e 1 4

d e s i r e d 33, 500 l b s . , ob ta ined from p re l imina ry s i z i n g . The d i f f e r e n c e is due t o :

1. 2 ,000 l b s of amno a r e included. 2. 3,265 l b s because of t h e much more f a v o r a b l e

engine weight (9 ,265- 6 ,000 ) . 3. t h e remaining -585 l b s is due t o round-off e r r o r s

i n t h e weight f r a c t i o n s .

The -585 l b s is d i s t r i b u t e d over a l l items which a r e computed wi th t h e weight f r a c t i o n s . T h i s d i s t r i b u t i o n is done i n p ropor t ion t o t h e i r component weight va lues i n t h e f i r s t column.

For example, t h e wing ad jus tment number is a r r i v e d a t by mul t ip ly ing -585 l b s by 6 ,922125 ,2510 .

Note:

The component weight v a l u e s i n t h e l a s t column a r e t h o s e t o be used i n t h e Class I weight and ba lance a n a l y s i s of t h e E r i s . This corresponds t o S t ep 1 0 as o u t l i n e d i n Chapter 2 , P a r t 11. The Class I weight and ba lance a n a l y s i s of t h e E r i s is c a r r i e d ou t i n Chapter 1 0 of P a r t I1 (See pp. 254-258) .

6: To save space , t h i s step is omit ted.


Part V Chapter 2 Page 1 6

3. CLASS I METHOD FOR ESTIMATING AIRPLANE INERTIAS P ~ P ~ ~ P P I I P P ~ ~ P ~ ~ = L ~ P P = ~ ~ P E ~ D ~ U . D . C P P P E = ~ % = ~ ~ ~ S = ~ = = ~ = = =

The purpose of t h i s chapter is t o provide a methodology f o r r ap id ly est imating a i r p l a n e i n e r t i a s . The emphasis is on rapid and on spending a s few engineering manhours a s poss ib le . Methods which f i t meet t h e s e ob jec t ives a r e r e fe r red t o as Class I methods. They a r e used i n conjunction w i t h t h e f i r s t s t a g e i n t h e prel iminary design process , t h e one r e f e r r e d t o as 'pod. sequence I ' i n P a r t I1 (Ref.2).

Sec t ion 3.1 p r e s e n t s a Class I method f o r est imating I and I,=. Ixx, yy These i n e r t i a moments a r e use fu l when-

ever it is necessary t o e v a l u a t e undamped n a t u r a l f r e- quencies and/or motion t i m e cons tan t s f o r a i r p l a n e s du- r ing pod. sequence I.

Example app l i ca t ions a r e discussed i n Sect ion 3.2.

The Class I method f o r a i r p l a n e i n e r t i a e s t ima t ion relies on t h e assumption, t h a t wi th in each a i r p l a n e category it is poss ib le t o i d e n t i f y a rad ius of gyra t ion , Rx,y#z f o r t h e a i rp lane . The moments of i n e r t i a of t h e

a i r p l a n e a r e then found from t h e following equations: 2

I = (Rx) Wlg (3.1)

Research i n References 9 , 1 0 and 11 has shown t h a t a non-dimensional r ad ius of gyra t ion can be associa ted w i t h each R component i n t h e following manner:

- R, = 2R,/e. with: e - (b + L) /2

The q u a n t i t i e s b and L i n E q n ~ ~ ( 3 . 4 ) and (3.5A are t h e wing span and t h e o v e r a l l a i r p l a n e length

-

respect ive ly .


Airp lanes of t h e same mission o r i e n t a t i o n tend t o have similar va lues f o r t h e non- dinensional r a d i u s of gy- r a t i o n . Tables B. 1 through B.12 (See Appendix B) p r e s e n t numerical v a l u e s f o r t h e s e non-dimensional r a d i i of gyratTon f o r d i f f e r e n t t y p e s of a i r p l a n e s .

The procedure f o r e s t ima t ing i n e r t i a s t h e r e f o r e b o i l s down t o t h e fo l lowing s imple s t e p s :

e~ 1: L i s t t h e v a l u e s of WTO, We, b, L and

e fo r t h e a i r p l a n e being designed.

I d e n t i f y which type of a i r p l a n e i n Tables B . l through B.12 b e s t ' f it 8 t h e a i r p l a n e being designed.

s e ~ 3: S e l e c t va lues f o r t h e non-dimensional r a d i i of g y r a t i o n corresponding t o WTO and WE. I t

must be kep t i n mind t h a t t h e d i s t r i b u t i o n of t h e mass d i f f e r e n c e between WTO and WE

is more important t h a n t h e mass d i f f e r e n c e i t s e l f .

Acquiring t h e knowledge of what t h e a i r p l a n e s i n Tables B . l through B.12 a r e l i k e is t h e r e f o r e e s s e n t i a l . A s u sua l , J a n e ' s (Ref. 8 ) is t h e sou rce f o r acqu i r ing t h a t knowledge.

Compute t h e a i r p l a n e moments of i n e r t i a from:

Values f o r b and f o r L fo l low from t h e a i r p l a n e threeview. The va lue f o r e fo l lows from Eqn. (3.6).

The reader w i l l have noted t h a t t h e r e is no r a p i d method f o r e v a l u a t i n g I,,. Thi s p roduc t of i n e r t i a can

be r e a l i s t i c a l l y eva lua ted on ly from a C l a s s I1 weight and ba lance a n a l y s i s . Such an a n a l y s i s is p re sen ted i n Chapter 9. I n t h e f i r s t s t a g e s of p re l iminary des ign I X Z

is no t u s u a l l y important . Therefore , it is normally


ignored u n t i l l a t e r s t a g e s i n t h e des ign process .

W p 5: Compare t h e es t imated i n e r t i a s of S t ep 4 wi th t h e d a t a of F i g u r e s 3.1 through 3.3. If t h e comparison is poor, f i n d an explana- t i o n and/or make adjustments .

B ~ D 6: Document t h e results ob ta ined i n S t e p s 1 through 5 i n a b r i e f , d e s c r i p t i v e r epo r t . I nc lude i l l u s t r a t i o n s where necessary.

Three example a p p l i c a t i o n s w i l l now be d i scussed :

3.2.1 Twin Engine P r o p e l l e r Driven Airplane: Se l ene 3 .2 .2 J e t Transpor t : Ourania 3 .2 .3 F i g h t e r : E r i s

3 .2 .1 Twin EnQiDe P r o u e r D-

S f e D l: The fol lowing in format ion is a v a i l a b l e f o r t h e Se lene a i r p l a n e :

WTO = 7,900 l b s WE 4,900 l b s b u 3 7 . 1 it

L = 43.0 it e = 40.05 f t ( P a r t 11, p.247, p.297)

SteD 2: From Table B3 (Appendix B ) t h e fol lowing a i r p l a n e s a r e judged t o be comparable t o t h e Se lene i n terms of mass d i s t r i b u t i o n : Beech D18S, Cessna 404 and Cessna 441.

S t e ~ 3: From Table B3 (Appendix B ) it is es t imated t h a t t h e fol lowing non-dimensional r a d i i of g y r a t i o n app ly t o t h e Selene:

S f s ~ 4 : With Eqns.(3.7) through (3 .9 ) t h e fol lowing moments of i n e r t i a can now be c a l c u l a t e d :


A t WE:

=xx - ( 4 . 9 0 0 / 7 . 9 0 0 ) ~ 7 . 5 9 8 - 4.713 s l u g f t 2 -

i~~ - (4 .900 /7 .900 )x l3 .109 = 8.131 s l u g f t 2

IZz - (4 .90017.900)x15.741 - 9.763 s l u g f t 2

S t e ~ 5: F i g u r e s 3 . 1 t h r o u g h 3.3 show t h a t t h e iner t ia estimates of S t e p 4 are r e a s o n a b l e .

Sf ;e~ 6: T h i s step has b e e n omitted t o save space.

S t e ~ 1: T h e following information is a v a i l a b l e fo r t h e O u r a n i a airplane:

w ~ o = 127.000 l b s W~ = 68 ,450 l b s b = 113 .8 f t

L = 127.0 f t e = 120.4 f t ( P a r t 11, p.251, p .299)

a e ~ 2: From T a b l e B7a (Appendix B) t h e f o l l o w i n g airplanes are judged t o b e c o m p a r a b l e t o t h e O u r a n i a i n terms o f mass d i s t r i b u t i o n : C o n v a i r 880, C o n v a i r 990, Boe ing 737-200, McDonnell Douglas DC8.

a e ~ 3: From T a b l e B7a (Appendix B) it is e s t i m a t e d t h a t t h e following n o n - d i m e n s i o n a l r a d i i of g y r a t i o n apply t o t h e O u r a n i a :

@D 4: W i t h Eqns. ( 3 . 7 ) t h r o u g h (3 .9 ) t h e f o l l o w i n g moments of i n e r t i a can now b e c a l c u l a t e d :

P a r t V C h a p t e r 3 Page 20

UeD 5: Comparison with Figures 3.1 through 3.3 i n d i c a t e s t h a t t h e i n e r t i a estimates of S tep 4 are reasonable.

ex> 6 : To save space, t h i s s t e p has been omitted.

3.2.3 F i a u

Stex> 1: The following information is a v a i l a b l e f o r t h e E r i s a i rp lane :

w ~ o = 64.905 l b s WE = 33,500 l b s b = 68.7 f t

L = 50.7 f t e = 59.7 f t ( P a r t 11, p.255, p.301)

e D 2 : From Table B9a (Appendix B ) t h e following a i r p l a n e s a r e judged t o be comparable t o t h e Er is i n terms of mass d i s t r i b u t i o n : DH Vampire 20 and Glos ter Meteor 11. The reader should no te t h a t t h e Vampire is t h e only jet f i g h t e r i n Table B9a with a twin boom configurat ion.

From Table B9a (Appendix B ) it is estimated t h a t t h e following non-dimensional r a d i i of gyra t ion apply t o t h e E r i s :

With Eqns. (3.7) through (3.9) t h e following moments of i n e r t i a can now be ca lcu la ted :


S f e ~ 5 : Comparison o f t h e r e s u l t s o f S tep 4 wi th Figures 3 . 1 through 3 . 3 i n d i c a t e t h a t t h e i n e r t i a e s t imate s a r e reasonable.

s e ~ 6 : This s t e p has been omitted t o save space .


PI127 . ..

3.2 C o a t i o n of P-ts of Inertia uLh Welaht

18

16

16

18 10' 10- - lo'

w e 3 . 3 Correlation of Y- M o m t s of Iner t ia with Weiaht

Page 23 Part v Chapter 3

4. CLASS I1 METHOD FOR ESTIMATING AIRPLANE COMPONENT = I S P P P I P I I P S P P I I P I = P ~ P ~ = ~ ~ ~ = O P ~ i l P n = P P ~ ~ ~ P ~ P = a = ~ ~ n S P =

WEIGHTS PeP=t==

The purpose of t h i s chapter is t o p resen t a Class I1 method f o r est imating a i r p l a n e component weights. Class I1 methods a r e those used i n conjunct ion with prel iminary design sequence I1 a s def ined i n P a r t 11, pp 18-23. The Class I1 weight est imating method accounts f o r such d e t a i l s a s :

1. Airplane take-off g ross weight

2. Wing and empennage design parameters such as :

a. a r e a b. sweep angle , c. t a p e r r a t i o , 5 d. th ickness r a t i o , t l c

3. Load f a c t o r , nlim o r nult

4. Design c r u i s e andlor d ive speed. VC o r VD

Note: i tems 3 and 4 fol low from a V-n diagram.

5. Fuselage conf igura t ion and i n t e r i o r requirements

6. Powerplant i n s t a l l a t i o n

7. Landing gear design and d i s p o s i t i o n

8. Systems requirements

9. Preliminary s t r u c t u r a l arrangement

To apply t h e Class I1 method f o r est imating component weights r equ i res a f a i r l y comprehensive knowledge about t h e a i r p l a n e being designed. T h i s knowledge was developed as a r e s u l t of pod. sequence I , discussed i n P a r t 11, pp 11-18.

Almost a l l a i r f rame manufacturers have developed t h e i r own Class I1 methods f o r es t imat ing a i r p l a n e component weights. Many of t h e s e methods a r e propr ie tary . The Class I1 methods used i n t h i s t e x t a r e based od t hose of References 1 2 , 1 3 and 14 . These methods employ empi- r i c a l equat ions which r e l a t e component weights t o a i r - plane design c h a r a c t e r i s t i c s such a s items 1-9 above.


The fol lowing b a s i c weight d e f i n i t i o n from P a r t I ( ~ ~ n l 2 . 1 7 ) w i l l b e used:

- =l

W~~ * + 'F + 'PL + 'tfo + W ~ r e ~ * ( 4 . 1 )

where: WE = empty weight, de f ined by Eqn.(4.2).

WF = mission f u e l weight , de f ined by: Eqn. ( 2 . 1 5 ) i n P a r t I.

WpL - payload weight, de f ined by t h e miss ion s p e c i f i c a t i o n and on page 8, P a r t I.

W t f o = weight of t rapped f u e l and o i l , found from p.7, P a r t I.

= crew weight, de f ined by t h e miss ion s p e c i f i c a t i o n and on page 8, P a r t I.

The C la s s I1 weight e s t ima t ing method t o be developed h e r e w i l l f ocus on e s t i m a t i n g t h e components of empty weight, WE which a r e de f ined a s :

where: = s t r u c t u r e weight , d i s cus sed i n Wstruct Chapter 5.

W Pwr

= powerplant weight , d i s cus sed i n Chapter 6.

'feq = f i x e d equipment weight, d i s cus sed i n Chapter 7.

I n Chapters 5-7 t h e s p e c i f i c C l a s s I1 methods a r e i d e n t i f i e d as fo l lows:

1. Cessna method: from Ref.12 2. USAF method from Ref. 13 3. GD (General Dynamics) method from Ref.13 4. Torenbeek method from Ref. 1 4

S e c t i o n 4 .1 p r e s e n t s a step- by- step procedure f o r us ing t h e C l a s s I1 weight e s t i m a t i o n method.

Sec t ion 4.2 p r e s e n t s a method f o r c o n s t r u c t i n g t h e V-n diagram, needed i n s e v e r a l of t h e weight equa t ions employed i n Chapters 5-7.

Example a p p l i c a t i o n s a r e p resen ted i n Sec t ion 4.3.


4 . 1 A METHOD FOR E S S T w m - I n t h i s s e c t i o n a step- by- step procedure is presen-

t e d f o r e s t ima t ing a i r p l a n e component weights and use t h e s e weights i n e s t ima t ing a i r p l a n e empty weight, WE.

A s w i l l be seen , t h i s p rocedure is i t e r a t i v e . The reason is, t h a t almost a l l a i r p l a n e component weights themselves are a f u n c t i o n of WTO. A f i r s t estimate f o r

WTo was ob ta ined dur ing t h e p re l imina ry s i z i n g of t h e

a i r p l a n e . The reader w i l l have n o t i c e d t h a t dur ing t h e C la s s I weight estimates (Chapter 21 , t h e o r i g i n a l est i- mate of WTO remained una l te red . Tha t w i l l no l onge r be

t h e case i n t h e Class I1 method.

The method is presen ted as p a r t of S t ep 2 1 i n pod . sequence 11, a s o u t l i n e d on p.19 of P a r t 11.

For t h e inexper ienced r eade r , it is suggested t h a t t h e fol lowing procedure be followed e x a c t l y as suggested.

WD 1: L i s t a l l a i r p l a n e components f o r which t h e weights are a l r e a d y known and t a b u l a t e t h e i r weights. Th i s in format ion can normally be ob ta ined from t h e miss ion s p e c i f i c a t i o n .

Typica l items of known weight are:

1. Payload 2. Crew 3. C e r t a i n o p e r a t i o n a l systems 4. C e r t a i n m i l i t a r y l oads 5. Engines ( t h e s e a r e sometimes s p e c i f i e d )

SteD 2: L i s t a l l a i r p l a n e components f o r which t h e weights w i l l have t o be es t imated . T h i s l ist w i l l c o n t a i n a t least t h e same items used i n Class I. However, p a r t i c u l a r l y i n t h e systems a r e a t h e l ist w i l l con ta in much more d e t a i l a t t h i s p o i n t .

I n p repa r ing t h i s l ist , use t h e groupings of components as i n d i c a t e d by Eqn.(4.2). Sub- d i v i s i o n of t h e s e groupings should be'done i n accordance wi th Chapte rs 5-7, Eqns. ( 5 . 1 1 , ( 6 . 1 ) and ( 7 . 1 ) .


- S f e ~ 3: Refer t o t h e s t r u c t u r a l arrangement drawing prepared under S tep 1 9 , p.19, P a r t 11.

- =s

The i n i t i a l s t r u c t u r a l arrangement drawing is needed t o i d e n t i f y those a r e a s of t h e s t r u c t u r e where s p e c i a l p rov i s ions were made o r where, because of a c l e v e r s t r u c t u r a l arrangement a weight saving can be claimed.

S t e ~ 4 : Determine from t h e t a b u l a t i o n below which . weight es t imat ion category b e s t r ep resen t s

t h e a i r p l a n e being designed.

1. Homebuilts 2. S ing le Engine Props 3. Twin Engine Props 4. Agr icu l tu ra l 5. Business J e t s 6. Regional Turboprops

below 1 2 , 5 0 0 l b s above 1 2 , 5 0 0 l b s

7. J e t Transports 8. M i l i t a r y Tra ine r s

low speed high speed

9. F igh te r s l o . M i l i t a r y P a t r o l , Bomb

and Transport Airplanes 11. Flying boats , Amphibi-

ous and F l o a t Airplanes small and low speed l a r g e and high speed

1 2 . Supersonic c r u i s e commercial

F igh te r and Attack P a t r o l , Bomb, Transp.

General Aviation Airplanes General Aviation Airplanes General Aviat ion Airplanes General Aviation Airplanes Commercial Transports

General Aviation Airplanes Commercial Transports Commercial Transports

General Aviation Airplanes F igh te r and Attack Airplanes F igh te r and Attack Airplanes M i l i t a r y P a t r o l , Bomb and Transport Airplanes

General Aviation Airplanes Commercial Transports and/or Mil.Patr . , Bomb and Transp.

Commercial Transports , but u s e F igh te r i n l e t d a t a F igh te r and Attack Mil.Patr . , Bomb and Transp.

The weight es t imat ion equat ions i n Chapters 5-7 a r e a l l given i n terms of t h e c a t e g o r i e s on t h e r i g h t s i d e of t h e above table.

a e ~ 5: Determine which equat ions i n Chapters 5-7 apply t o t h e a i r p l a n e f o r which t h e Class I1 weight e s t ima te is t o be made. L i s t t h e s e equat ions f o r each weight component.


S t e ~ 6: Make a list of a l l required input d a t a needed i n t h e equat ions of S tep 5.

-D 7 : Compute t h e component weights with t h e appl icable equat ions of S tep 5.

Notes:

1. The reader w i l l observe t h a t Chapters 5-7 o f t e n contain more than one equat ion t o es t imate t h e weight of a p a r t i c u l a r component. In t h a t case e s t ima te t h e weights with a l l appl icable equat i- ons U e an ave-

2. Sometimes it is d e s i r a b l e t o ' c a l i b r a t e ' a component weight equat ion with t h e he lp of known weight d a t a from e x i s t i n g a i rp lanes . The component weight d a t a of Appendix A can be used f o r t h i s purpose. C a l i b r a t i o n is done by applying t h e weight equat ions t o t h e appropr ia te compo- nen t s and comparing t h e answers with t h e a c t u a l weight d a t a of Appendix A. The so- cal led 'fudge- cons tan t s ' which appear i n a l l Class I1 weight equat ions can then be a l t e r e d t o obta in a b e t t e r es t imate. The reader should be c a r e f u l and only use t h i s ' c a l i b r a t i o n ' method i n conjunction with a i r p l a n e s which have s i m i l a r missions.

3 . I n t h e systems area , t h e r e a r e n o t enough r e l i - a b l e equat ions ava i l ab le . I n t h a t case it is de- s i r a b l e t o a l s o es t ima te t h e average app l i cab le weight f r a c t i o n f o r each system component. Th i s can be done w i t h t h e d a t a i n Appendix A. The examples i n Sect ion 4.3 show how t h i s is done.

Add a l l component weights and obta in an es t imate f o r WE, from Eqn. (4 .2 ) .

S f e ~ 9: Compute a new value f o r WTO with Eqn.(4.1),

but : 1. use f o r WE t h e va lue obtained i n S tep 8.

2. use f o r Wp a value obtained from

Eqn. (2 .15 ) i n P a r t I. This imlies t h a t t h e mission f u e l needed m u s t be adjus ted f o r t h e new value of WTO. The r e s u l t is:


= ('E + 'PL + 'crew ) / t n f f ( i + M,,,) - M,,, - M t f o l

i a l u e s f o r M t f D Mres and %fo were a l r e a d y ob ta ined

dur ing t h e p re l imina ry s i z i n g work desc r ibed i n Chapter 2 of P a r t I. These f r a c t i o n s may have changed i f , du r ing t h e C l a s s I drag p o l a r a n a l y s i s of Chapter 1 2 , P a r t I1 a s i g n i f i c a n t change i n L I D was discovered. I n t h a t c a s e it was recommended i n S t ep 1 4 , P a r t I1 (p.16-17) t o redo t h e p re l imina ry s i z i n g . T h i s i n t u r n would r e s u l t i n new v a l u e s f o r t h e f r a c t i o n s i n Eqn.(4.3).

e~ 10: Use t h i s new estimate f o r WTO t o i t e r a t e

back through S t e p s 7-9 u n t i l t h e WTO v a l u e s ag ree w i t h i n 0.5 percen t .

1. If t h e new v a l u e of WTO ob ta ined i n S t ep 9 d i f -

f e r s from t h e o r i g i n a l one by more t h a n 5 p e r c e n t it w i l l b e necessary t o account f o r t h e e f f e c t on r equ i r ed engine t h r u s t ( o r power) a t take- off . Th i s i n t u r n w i l l a f f e c t t h e engine weight.

2. Accounting f o r a change i n r equ i r ed take- off t h r u s t ( o r power) may be done by us ing t h e r a t i o (TIWITO ( o r (W/P)TO obta ined from t h e p re l imina ry

s i z i n g p roces s of Chapter 3 , P a r t I.

e D 11: Document a l l c a l c u l a t i o n s i nc lud ing a l l assumptions made, a l l d e c i s i o n s made and a l l i n t e r p r e t a t i o n s made i n a b r i e f , d e s c r i p t i v e r e p o r t . Where needed, i n c l u d e c l e a r l y drawn ske tches .

Inc lude a f i n a l C l a s s I1 weight s t a t emen t , us ing t h e groupings sugges ted by Eqns. ( 4 . 2 1 , ( 5 . 1 1 , ( 6 . 1 ) and ( 7 . 1 ) .


4.2 METHODS FOR CO-G V - n DI;AGRAMS

I n t h i s s e c t i o n a step- by- step procedure is presen ted f o r c o n s t r u c t i n g V-n diagrams f o r t h e fol lowing t y p e s of a i r p l a n e s :

4 . 2 . 1 FAR 2 3 C e r t i f i e d Ai rp lanes 4 .2 .2 FAR 2 5 C e r t i f i e d Ai rp lanes 4.2.3 M i l i t a r y Ai rp lanes

Example a p p l i c a t i o n s f o r t h r e e a i r p l a n e s a r e provided i n sub- sect ion 4.2.4.

The V-n diagrams a r e used t o de te rmine des ign l i m i t and des ign u l t i m a t e load f a c t o r s as w e l l as t h e corresponding speeds t o which a i r p l a n e s t r u c t u r e s a r e designed. A s w i l l be seen i n t h e Class I1 weight equa t ions of Chapters 5-7, many r e q u i r e a s i npu t a des ign load f a c t o r and /or a des ign speed.

no t e s :

1. The V-n diagrams g iven h e r e a r e s i n p l i f i e d v e r s i o n s of t h o s e def ined i n Refs 15- 17. They should be used on ly i n conjunc t ion with C l a s s I1 weight e s t i m a t i o n methods.

2. I n t h e C l a s s I1 method on ly f l aps- up cases a r e considered.

2 3 C w d A i w . 4.2.1 V-n Di-am f o r FAR

Reference 1 5 , i n P a r t 23.335 p r e s e n t s t h e V-n diagram shown i n F igu re 4.1. The fo l lowing d e f i n i t i o n s apply t o t h e v a r i o u s speeds g iven i n t h e diagram:

Note: a l l speeds are normally g iven i n KEAS.

Vs = +Ig s t a l l speed o r t h e mininum speed a t which t h e a i r p l a n e is c o n t r o l l a b l e

VC = des ign c r u i s i n g speed

VD = des ign d iv ing speed

VA - des ign maneuvering speed

Determinat ion of t h e s e speeds and d e t e r m i n a t i ~ of t h e c r i t i c a l p o i n t s A, C , D , E , F and G i n ~ i g u r e 4.1 is d i scussed i n sub- sub- sections 4.2.1.1 through 4.2.1.7.


e 4.1 V-n D i m c o r d j . p a t o FAR 2 3

where: GW = f l i g h t des ign g r o s s weight i n l b s

S - wing a r e a i n f t 2

p = a i r d e n s i t y i n s l u g s l f t 3

C = maximum normal f o r c e c o e f f i c i e n t . Nmax

The maximum normal f o r c e c o e f f i c i e n t fo l lows from:

l 2 + (C ) 2 } 1 1 2 C~ = { ( C L max max Dat C, JA max

I n p re l imina ry des ign it is accep tab le t o set:

C~ = 1 . 1 C max Lmax

where t h e c o n s t a n t k c t a k e s on t h e fol lowing va lues :

kc - 3 3 f o r normal and u t i l i t y ca t ego ry a i r p l a n e s up t o WIS = 2 0 p s i .


kc v a r i e s l i n e a r l y from 33 t o 28.6 as WIS v a r i e s from 20 t o 100, f o r normal and u t i l i t y ca tegory a i r p l a n e s .

k c = 3 6 f o r a c r o b a t i c ca tegory a i r p l a n e s .

U te : Vc need n o t be more than 0.9VH, where VH

is t h e maximum l e v e l speed ob ta ined wi th maximm power o r wi th maximum t h r u s t .

. . 4.2.1.3 D e t e r w o n of &&g~ dl-a s ~ e d . VD

VD ( o r RID)> 1.25VC (o r 1.25Mc). ( 4 . 8 )

where: VC fo l lows from Eqn.(4.7).

where: nlim is t h e l i m i t maneuvering load f a c t o r g iven

by Eqn. ( 4 . 1 3 ) .

Note: VA need n o t exceed VC

:e C , is g iven by: w h e ~ max

neg 2 )2 ]112

= ) + ( C D . _ (4 .11) C , n max neg

U max neg

I n p re l imina ry des ign it is accep tab le t o use:

C~ = 1 . 1 C max Lmax

ne9 neg where: C, is t h e maximum nega t ive lift c o e f f i c i e n t .

Y max neg


1.2.1.6 of loan factor. n . . l i m

The p o s i t i v e , des ign l i m i t l oad f a c t o r is g iven by:

" l i m need no t be g r e a t e r t han 3.8

" l i m = 4.4 f o r u t i l i t y ca t ego ry airplanes

" l i m = 6.0 f o r a c r o b a t i c a i r p l a n e s

The nega t ive , de s ign l i m i t load f a c t o r is g iven by:

"lim 2, 0*4nlim f o r normal and f o r u t i l i t y (4.14) neg ca t ego ry airplanes

>, 0. 5nlim f o r a c r o b a t i c airplanes (4.15)

4.2.-1.7 Cons t ruc t ion o f w t l oad i m o r lines Fia .4 . l

The g u s t load f a c t o r l i n e s i n F i g u r e 4.1 are def ined by t h e fol lowing equa t ion :

where: K is t h e g u s t a l l e v i a t i o n f a c t o r g iven by: 9

Kg = 0. 88pg/(5.3 + p 1 ,

g (4.17)

where:

The de r ived g u s t v e l o c i t y , Ude is def ined as fo l lows:

- cmt uJles: 'de = 50 f p s between sealevel and 20,000 f t

'de = 66.67 - 0.000833h between 20,000 and 50,000 f t

Forthe u u s t lines:

'de = 25 f p s between sealevel and 20,000 f t



Reference 1 6 , i n P a r t 25.335 p r e s e n t s t h e two V-n diagrams shown i n F igu res 4.2a and 4.2b. The fol lowing d e f i n i t i o n s apply t o t h e v a r i o u s speeds g iven i n t h e diagrams:

Note: a l l speeds a r e normally g iven i n KEAS.

= +1-g s t a l l speed o r t h e minimm s t e a d y f l i g h t "1 speed which can be ob ta ined

VC = des ign c r u i s i n g speed

VD = des ign d iv ing speed

VA = des ign maneuvering speed

VB - des ign speed f o r maximum g u s t i n t e n s i t y

Determinat ion of t h e s e speeds and de te rmina t ion of t h e c r i t i c a l p o i n t s A, D, En F D H, B', C ' , D ' , E ' , F ' and G' is d iscussed i n sub- sub- sections 4.2.2.1 through 4.2.2. 8.

where: GW = f l i g h t des ign g r o s s weight i n l b s

S = wing a r e a i n f t 2

P = a i r d e n s i t y i n s l u g s l f t 3

C = maximum normal f o r c e c o e f f i c i e n t , as Nmax computed from Eqn.(4.5) o r ( 4 . 6 ) .

VC must be s u f f i c i e n t l y g r e a t e r t han VB t o p rov ide

f o r i n a d v e r t e n t speed i n c r e a s e s l i k e l y t o occur as a r e s u l t of s e v e r e a tmospher ic tu rbu lence . For VB, see sub- sub- section 4.2.2.5.

+ 43 k t s


w e 4 . 2 a V-n --am Ac-a t o FAR 2 5

VD S .PE ED,V - KEAS


VD (o r MD) > l.25Vc (o r 1.25MC)

where: VC fol lows from Eqn.(4.20).

where: nlim is t h e l i m i t maneuvering load f a c t o r a t VC.

The l i m i t maneuvering load f a c t o r i n Eqn. (4.22) follows from 4.2.2.7 or from 4.2.2.8 depending on which is t h e more c r i t i c a l .

VA need n o t exceed VC.

4.2.2.5 Dete-on of -an s ~ e e d f o r maxirmm a m

VB need not be g r e a t e r than VC.

VB may n o t be less than t h e speed determined from

t h e i n t e r s e c t i o n of t h e CN l i n e and t h e g u s t l i n e marked VB. max

The negat ive s t a l l speed l i n e i n F igure 4.2a is determined with t h e method of sub-sub-section 4.2.1.5.

4.2.2.7 Dete-on of de- f a c t o r . nlim

The p o s i t i v e l i m i t maneuvering load f a c t o r , n l i % o s

is determined from:

n Z 2.5 a t a l l times l i%0s4 *

"limpos need no t be great -er than 3 .8 a t WTc

The negat ive, design l i m i t load f a c t o r is determined from:


a l i m n - varies l i n e a r l y from t h e va lue a t VC t o

zero a t VD

4.2.2.8 C o w t i o n of aUQf l oad f a c t o r lines in Fia.4.2h

The g u s t l oad f a c t o r l i n e s i n F i g u r e 4.2b are a r r i v e d a t w i t h t h e h e l p of Eqns. (4.16) th rough (4.18). The d e r i v e d g u s t ve loc i t ies , Ude i n FAR 25 are as fo l l ows :

For t h e a u s t l i n e &ed VB:



Ude = SO f p s between sealevel and 20,000 f t


For t h e uust m e m e d b:



4.2.3 V-n D i a w f o r A i w . .

Reference 1 7 , p rov ides t h e V-n d iagram g i v e n i n F i g u r e 4.3. The i n d i c a t e d l i m i t l o a d factors must no t b e less t h a n t h o s e d e f i n e d i n Tab le 4.1.

The speeds i n F i g u r e 4.3 are normal ly g i v e n i n KEAS and are d e f i n e d as fo l l ows :

VH = maximum l e v e l speed which can be a t t a i n e d a t t h e combination of weight and a l t i t u d e under c o n s i d e r a t i o n

VL = maximum d i v e speed, t y p i c a l l y 1. 25VH

G u s t l i n e s are as i n FAR 25. G u s t induced load f a c t o r s are normal ly n o t c r i t i c a l f o r m i l i t a r y airplanes w i t h l i m i t l o ad f a c t o r s above 3.00.

P a r t v Chapter 4 Page 38

USN

7

I 1 I cL 0 I

VH

I

Page 39

"L

SPEFD,~ ~v KCAS

4 G R E F . \7

( ~ E G T A S C E 4.1)

e 4.3 V-n D i m A c c o r m a t o m - A 8 8 6 1 0

T a b l e 4.1 L i m i t Load F a c t o r s f o r M i l i t a r y Airplanes ~ P I P P I P P I P I I I E P P E I P ~ I P ~ ~ P . I ~ . C P P I E ~ ~ ~ I I ~ ~ ~ ~ ~ P ~ ~ ~ E ~ ~ ~ E

Airplane Type L i m i t Load F a c t o r , nlim

a t F l i g h t Des ign Gross Weight , GW

USAF P o s i t i v e Nega t ive

F i g h t e r 8.67 -3.00

A t t a c k F i g h t e r , A t t a c k , 7.33 -3.00 T r a i n e r

O b s e r v a t i o n 6.00 -3.00

T r a i n e r 5.67 -2.33

U t i l i t y U t i l i t y 4.00 -2.00

S m a l l Bomber 3.67 -1.67

Medium Bomber, P a t r o l , Weather , A s s a u l t Transp . Ant i- submar ine , 3.00 -1.00

Reconnaissance

Medium Transp. 2.50 -1.00 -t

Heavy Bomber, Heavy Transp . 2.00 -1.00

P a r t V C h a p t e r 4

T h e fol lowing example appl icat ions w i l l b e d i s c u s s e d :

-

4.2.4.1 Twin E n g i n e P r o p e l l e r Driven A i r p l a n e : S e l e n e

4.2.4.2 Jet T r a n s p o r t : O u r a n i a 4.2.4.3 F i g h t e r : E r i s

4.2.4.1 Twin w e P r o ~ e u e r Driven A i w

A c c o r d i n g t o t h e mission s p e c i f i c a t i o n ( T a b l e 2.17, P a r t I) t h i s is a FAR 23 airplane. It w i l l b e assumed t h a t u n d e r FAR 2 3 it w i l l b e c e r t i f i e d u n d e r t h e n o r m a l category.

S i n c e CL = 1 .7 ( P a r t I, p . 1 7 8 ) , it f o l l o w s max

from Eqn.(4 .6) t h a t : C = 1.1x1.7 = 1.87. Nmax

S i n c e (WIS)TO = 46 psf ( P a r t I , p .178) . t h e value

f o r s t a l l s p e e d as found from Eqn. ( 4 . 4 ) is1

Vs = ~ 2 ~ 4 6 1 0 . 0 0 2 3 7 8 ~ 1 . 8 7 ~ 'I2 = 1 4 4 f p s = 85 k t s .

The d e s i g n wing l o a d i n g f o r t h e S e l e n e is 46 ps i . T h i s y i e l d s kc = 31.6. With Eqn. ( 4 . 7 ) t h i s i n t u r n

g i v e s V 3 214 k t s . C

T h e S e l e n e was t o h a v e a c r u i s e speed of 250 k t s a t 75 percent power a t 1 0 , 0 0 0 f t ( P a r t I , T a b l e 2.17.). F o r 1 0 0 percent power t h i s would y i e l d a maximm cruise s p e e d

wh ich is a factor (100175) 'I3 = 1.1 h i g h e r , o r 275 k t s . A c c o r d i n g t o s u b- s u b- s e c t i o n 4.2.1.2, VC n e e d no t b e h i g h e r t han 0 . 9 ~ 2 7 5 = 2 4 8 k t s .

Thus: VC = 2 4 8 k t s .

A c c o r d i n g t o s u b- s u b- s e c t i o n 4.2.1.3, t h e d e s i g n d i v e s p e e d is: VD = 1 . 2 5 ~ 2 4 8 = 310 k t s .


The p o s i t i v e l i m i t load f a c t o r of t h e Selene as given by Eqn.(4.13) is:

The negat ive l i m i t load f a c t o r as given by Eqn. (4.14) is:

on of Gwt Load F

The o v e r a l l a i r p l a n e l i f t c u r v e s lope , C,. can Y

Q-

be shown t o be 0.095 deg-'= 5.44 rad-l. With c = 4.92 it (Table 13.1, P a r t 1 1 1 , t h e value of p is: 44.8, according t o Eqn. (4.18). 9

The g u s t a l l ev i a t i on f a c t o r fol lows from Eqn.(4.17) as :

The g u s t load f a c t o r l i n e s now fol low from Eqn. ( 4 . 1 6 ) as:

n - 1 + 0.0094V f o r t h e VC l i n e and: li$USt -

n = 1 + 0.0047V f o r t h e VD l i n e . l i m g u s t

on of \IA From Eqn. (4.9): VA 85x(3.44) 'I2 = 158 k t s .

It w i l l be assumed t h a t CL I - 1.18. This max

neg y i e l d s CN P - 1.3. Using Eqn.4.4 it is found t h a t

max neg

A7

t h e negat ive l g s t a l l speed is 102 k t s .

With t h e s e d a t a it is now p o s s i b l e t o draw t h e V-n diagram for t h e Selene. The result is shown i n Fig.4.4.


w e 4 . 5 a V - n Maneuver D i m I

1 I

3.0 -

200

w e 4.Sb E x w l e V - n Gust D a for t h e 0-


A c c o r d i n g t o t h e mission specif icat ion ( T a b l e 2 - 1 8 , P a r t I) t h i s is a FAR 25 airplane.

On Of Y *

S i n c e C = 1 . 4 ( P a r t I , p.1841, it follows Lmax

from E q n . ( 4 . 0 t h a t : CN = 1.1x1.4 = 1.54. max

S i n c e (W/S)TO = 9 8 psf ( P a r t I, p .184) . t h e va lue

f o r s t a l l speed as f o u n d f r o m Eqn.(4 .4) is:

vs = I 2 ~ 9 8 / 0 . 0 0 2 3 7 8 ~ 1 . 5 4 ) 'I2 = 2 3 1 f p s - 1 3 7 k t s .

VA f o l l o w s from t h e in te r sec t ion of t h e + l g s t a l l

l i n e and t h e +2.50 l o a d factor l i n e : VA = 1 9 5 k t s .

VB f o l l o w s from t h e intersect o f t h e + l g s t a l l l i n e

and t h e VB g u s t l i n e . T h i s intersect w i l l b e d e t e r m i n e d

upon calcula t ion of t h e VB g u s t l i n e .

Acco rd ing t o Eqn. (4 .20) : VC VB + 43 k t s .

T h e r e f o r e : VC 1 9 5 + 43 = 238 k t s . However, t h e

miss ion specification of T a b l e 2 .18 ( P a r t I ) ca l ls for a cruise s p e e d o f M = 0.82 a t 35 ,000 f t . T h i s c o r r e s p o n d s t o 483 k t s a t 35 ,000 f t or a dynamic pressure of 296 p s f . A t sealevel, t h e c o r r e s p o n d i n g v a l u e i n KEAS is 295 k t s . S i n c e t h i s is larger 2 3 8 k t s , VC = 295 k t s .

A c c o r d i n g t o s u b- s u b- s e c t i o n 4.2.2.3, t h e d e s i g n d i v e s p e e d is: VD = 1.25xVC = 1 . 2 5 ~ 2 9 5 = 369 k t s . ,

Determina t ionof l im

The pos i t i ve l i m i t l o a d factor o f t h e O u r a n i a as

P a r t V C h a p t e r 4 P a g e 43

g iven by Eqn.(4.23) is: -

%igos 2 . 1 +{24,000/(127,000 + 10,000)) = 2.28

The except ions i n sub-sub-section 4.2.2.7 demand t h a t t h i s load f a c t o r never be less than 2.5. Therefore:

The negat ive l i m i t load f a c t o r is -1 up t o VC and varies l i n e a r l y t o zero a t VD.

The o v e r a l l a i r p l a n e l i f t c u r v e s lope , CL can a-

be shown t o be 0.085 deg-'= 4.87 rad-'. With c = 1 2 . 5 f t (Table 13.2, P a r t 11). t h e va lue of r is: 42.0, according t o Eqn. (4.18). 4

The g u s t a l l e v i a t i o n f a c t o r fol lows from Eqn.(4.17) as :

The g u s t load f a c t o r l i n e s now fol low from Eqn. (4.16) as:

n = i + 0.0051V f o r t h e VB l i n e . l i m g u s t

n = 1 + 0.0039V f o r t h e VC l i n e and: lim,ust -

n = 1 + 0.0019V f o r t h e VD l i n e . l imgus t

From t h e i n t e r s e c t i o n of t h e +lg s t a l l l i n e with t h e VB g u s t l i n e it follows t h a t VB = 195 k t s .

I t w i l l be assumed t h a t C = - 1.00. This Lmax

n eg y i e l d s C. re - 1.1. Using Eqn.4.4 it is found t h a t

Nmax neg

t h e negat ive l g s t a l l speed is 162 k t s .


With t h e s e d a t a it is now p o s s i b l e t o draw t h e V-n diagram f o r t h e Selene. The resul t is shown i n F igu res 4.5a and 4.5b.

4.2.4.3 F i w

According t o t h e miss ion s p e c i f i c a t i o n (Table 2.19, P a r t I) t h e Se lene is an at tack f i g h t e r . From Table 4.1 it fo l lows t h a t nli = 7.33 and nlim = - 3.0.

mpos neg

The maximum l e v e l speed a t s e a l e v e l is VH = 450 k t s .

The des ign d i v e speed, VL = 1 . 2 5 ~ 4 5 0 = 563 k t s .

The g u s t l i n e s are f a r w i th in t h e maneuvering V-n diagram and are n o t computed for t h i s a i r p l a n e . Fig.4.6 p r e s e n t s t h e V-n diagram f o r t h e E r i s .

8 - 7.53 _ _ - - - -

6 -

-

4 - -

s

fVHirso *v I

4 00 SPEED, v- KEbS

vD.scam 600

- 2 - ?

*

-4 - -- - p a r t v Chapter 4 Page 45

4 . 3 L E APPLICATIONS FOR U S I1 WEIGHT ESTI-

I n t h i s sect ion, t h r e e example a p p l i c a t i o n s of t h e Class-I1 weight es t imat ion method desc r ibed i n Sec t ion 4 . 1 are d i scussed :

=s

4.3.1 Twin Engine P r o p e l l e r Driven Airplane: Se l ene 4.3.2 Jet Transpor t : Ourania 4.3.3 F igh te r : E r i s

4.3.1 Twin E n a h e P r o p

S t e ~ 1: The fol lowing weight items are a l r e a d y known :

From Table 10.4, P a r t 11:

Payload : w~~ = 1,250 l b s

Fuel : WF = 1,706 l b s

TFO : - W t f o

- 44 l b s

From P a r t 11, p.135:

Engine d ry weight: We = 1,400 l b s

a e ~ 2: Weights need t o be es t imated f o r t h e fol lowing items:

W-struct:

1) Wing 2 ) Adjustment f o r Fowler f l a p s 3) Empennage

4 ) Fuselage 5) Nacelles 6) Landing Gear

1) Engines 2) A i r i nduc t ion system 3) P r o p e l l e r s

4 ) Fuel System 5) Propuls ion i n s t a l l a t i o n

1) F l i g h t c o n t r o l s 2 ) Electrical system

3 ) Ins t rumenta t ion , a v i o n i c s and e l e c t r o n i c s

4 ) ~ i r - c o n d i t i o n i n g and de- icing 5 ) Oxygen

6) Furn ish ings 7 ) p a i n t


S f e ~ 3: The s t ruc tu ra l arrangement drawing f o r t h e Se lene is p re sen ted i n Chapter 8 of P a r t 111.

S t e D 4: From a weight e s t ima t ing viewpoint t h i s a i r p l a n e f a l l s i n t h e General Avia t ion Ai rp lane category.

SteD The fol lowing weight equa t ions apply t o t h e Se lene :

'struct : 1) Wing: Eqns (5.4) and (5 .5)

2 ) Adjustment f o r Fowler f l a p s : an ex t ra f a c t o r of 2 pe rcen t w i l l be added i n accordance wi th 5.2.2.2.

3 ) Empennage: Eqns (5.14) - (5.16)

4 ) Fuselage: Eqns (5.25) and (5.27)

5 ) Nacelles: Eqn. (5.33)

6) Landing Gear: Eqns (5.40) and (5.42)

wpwr : 1) Engines: see S tep 1.

2) A i r i nduc t ion system: Eqn.(6.8)

3 ) P r o p e l l e r s : Eqns (6.13) and (6.14)

4 ) Fuel System: Eqns (6.17) and (6.18)

5 ) Propuls ion system: Eqns (6.3) and (6 .4)

'feq: 1) F l i g h t c o n t r o l system: Eqns ( 7 . 1 1 , (7 .2) and (7 .4) . Note: h y d r a u l i c s and pneumatics are included i n i t e m 1).

2 ) Electrical system: Eqns (7.12) - (7.14)

3 ) Ins t rumenta t ion , avionics and electro- n i c s : Eqn. (7.21)

4 ) Air- condi t ioning + de- ic ing : Eqn. (7.2 8)

5) Oxygen system: Eqn. (7.35) 37

6) Furn ish ings : Eqns (7.41) and (7.43)

7 ) P a i n t : Table A3.2a


SteD The following l ist i temizes a l l required input-da ta f o r es t imat ing t h e weight items l i s t e d in s t e p s 2 and 5.

-1

7 ,900 l b s nlim = 3 .44 S = 1 7 2 f t 2 W ~ O

VC = 2 4 8 k t s VD = 310 k t s nult = 5.16

A = 8 5 = 0.4 A 1 1 4 = 0 deg.

Notes: 1) The value f o r nlim fol lows from t h e V-n diagram of Figure 4.4.

2 ) Most d a t a were obtained from Selene d a t a l i s t e d i n P a r t 11. The reader is reminded t h a t a d e t a i l e d geometric d e f i n i t i o n may be found i n P a r t I1 a s Table 13.1, a Class I we igh t s ta tement a s Table 10 .4 . Detai led d e f i n i t i o n s of fuse lage , wing, t a i l s , landing gear and powerplant may be found i n Chapters 4 , 5 , 6 , 7 , 8 and 9 r e spec t ive ly i n P a r t 11.

S f e ~ 7: T a b l e 4.1 lists a l l weights computed a s p a r t of t h e Class I1 weight es t imat ion process. Observe t h a t t h e Class I weight es t imates (computed from weight f r a c t i o n s ) a r e averaged i n t o t h e new weight c a l c u l a t i o n s t o form t h e Class I1 weight est imate.

S t e ~ 8: The Class I1 empty weight of t h e Selene is 5 ,122 l b s . This compares with 4 ,900 l b s f o r t h e Class I weight est imate. This r ep resen t s a d i f f e r e n c e of 222 l b s which is 4.5 percent of t h e Class I empty weight.


Several comments a r e i n order :

1. an i t e r a t i o n through t h e equat ions of S tep 7 should be performed, t o determine t h e 'convergencep empty weight.

2. s eve ra l weight savings can be made i n t h e Selene:

a ) by manufacturing t h e p r o p e l l e r s out of composites, t h e i r weight can probably be c u t by 4 0 percent f o r a weight saving of 93 l b s .

b ) t h e empennage can be manufactured from composites which would y i e l d a weight saving of about 1 5 percent , o r 2 4 lbs .

c ) t h e n a c e l l e s can be manufactured p a r t i a l l y from composites which would y i e l d a weight saving of about 1 0 percent , o r 2 6 lbs .

d l by manufacturing t h e low s t r e s s a reas of t h e wing and fuse lage from composites, a weight saving of about 5 percent should be f e a s i b l e . This would save 72 lbs .

e ) combining a ) through d l y i e l d s a saving of 215 l b s . I t t h e r e f o r e appears q u i t e p o s s i b l e t o br ing t h e o v e r a l l Selene take-off weight i n a t t h e o r i g i n a l e s t ima te of 7 ,900 l b s .

9 and 10 : Not needed, s e e i tem e l , S tep 8.

S f i e ~ 11: To save space, t h i s s t e p has been omitted.

B€DE 5 5

P a r t v Chapter 4 Page 4 9

T a b l e 4 . l a C l a s s I1 W e i g h t E s t i m a t e s f o r t h e S e l e n e ...................................................

Compo-nent Methods : U s e as - C l a s s I USAF T o r e n b e e k C l a s s I1 * P a g e 9 E s t i m a t e

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . --------------- S t r u c t u r e w e i g h t , Ws t ruc t :

Wing 7 3 8 5 80 410 576

A d j u s t m e n t f o r F o w l e r f l aps , 2 p e r c e n t : 1 2

Empennage 1 7 9 1 4 9 1 5 5 1 6 1

F u s e l a g e 62 1 830 1 , 1 3 0 86 0

Nacelles 249 N. A. 272 2 6 1

L a n d i n g g e a r 3 80 1 9 6 3 1 3 296 ....................................................

Wstruct 2 , 1 6 7 1 , 7 5 5 2 , 2 8 0 2 , 1 6 6 ..................... e x c l . n a c . ===================

P o w e r p l a n t w e i g h t , W : ====================gEg=============================

E n g i n e s 1 , 4 0 0 1 , 4 0 0 1 , 4 0 0 1 , 4 0 0

A i r i n d u c t i o n i n pwrplt 8 8 8 8

P r o p e l l e r s 200 2 5 0 2 5 0 233

F u e l sys tem i n p w r p l t 157 1 3 5 1 4 6

P o w e r p l a n t i n s t . 1 0 8 1 0 8 = = = = 1 = = = = = = = = 1 5 = = = = = P = 1 3 1 = = = P l t t B P = = = t P = = = = = = = = = = = = =

W 1 , 7 0 8 2 , 3 1 9 2 , 3 0 0 1 , 9 7 5 =EN5====I===PI===========P1=P============P=131==============

i n c l u d e s eng ine a n d propeller w e i g h t , Eqn . (6 .3 )

' i n c l u d e s eng ine a n d propeller w e i g h t , Eqn. ( 6 . 4


T a b l e 4 . l b C l a s s I1 Weight E s t i m a t e s f o r t h e S e l e n e = D = I = 5 = P D P P = = = P S = P I = P P = t = = = = P = = = ~ = O I = . c P P n n = ~ = = ~ ~ = = n

Component Methods : Use as C l a s s I C e s s n a USAF T 'beek C l a s s I1 P a g e 9 E s t i m a t e

---= -,, ===%I=~=======P==PS===========~P==~=PI=P===~=======

F i x e d equipment w e i g h t , Wfeq:

P l = P l = P = = = P P = P P = = = S I '=P===P3P=PSP=IP=tP==P============

Wfc 1 3 3 294 9 1 1 7 3

Whps: t h i s is i n c l u d e d i n W f c

'els 212 210 209 210

1 0 3 1 0 3 'iae 88 88 'api

GD: 2 5 2 5

'fur 2 5 8 410 3 34

W T a b l e A3.2a: 4 8 lE~llpP=PD=PP=I=P=I=PP===r:===3:I===st5==P==xx=~~~=~==== W 1 , 0 2 5 not complete 9 81 =f S ~ = ' ~ = S ~ ~ = = P I = L = = = = = ~ = : = = L ~ = = = = = = = P = O I = = = = = = = = = = = = P = = = =

C l a s s I1 empty w e i g h t . WE fol lows f r o m Eqn . (2 .1 ) :

= 2.166 + 1 , 9 7 5 + 981 = 5 ,122 l b s

ARMSTRONG WH ITMRTH

ARGOSY 222

,

>

C h a p t e r 4 P a r t V P a g e 5 1

WD 1: The fol lowing weight items a r e a l r e a d y known : -

=z

From Table 10.5 , P a r t 11:

Payload weight: w~~ = 30 ,750 l b s

Crew weight: 'crew = 1 , 0 2 5 l b s

Fuel weight: w~ = 25,850 l b s

Trapped f u e l and o i l : W t f o = 9 2 5 l b s

From P a r t 11, p.138:

Engine d ry wht : = 9 , 2 2 4 l b s

e D 2 : Weights need t o be es t imated f o r t h e fo l lowing items:

S t r u c t u r a l W e i q b t . Wstruct :

1) Wing 2 ) Adjustment f o r Fowler f l a p s 3 ) Empennage

4 ) Fuselage 5) Nace l les 6 ) Landing Gear

power^-t . Wpwr :

1) Engines 2 ) Fuel system 3 ) Propuls ion system

4 ) Accessory d r i v e s , s t a r t i n g and i g n i t i o n system

5 ) Th rus t r e v e r s e r s

t ureigbt. Wfeq:

1) F l i g h t c o n t r o l s 2 ) E l e c t r i c a l system

3 Ins t rumenta t ion , a v i o n i c s and e l e c t r o n i c s

4 ) Ai r- condi t ion ing , p r e s s u r i z a t i o n and de- icing

5 ) Oxygen 6 ) APU 7 ) Furn ish ings 8) Baggage and

cargo handl ing 9 ) Opera t iona l i tems 1 0 ) P a i n t

a e ~ 3: The s t r u c t u r a l arrangement drawing f o r t h e Ourania is p re sen ted i n Chapter 8 of P a r t 111.


From a weight e s t ima t ing viewpoint t h i s a i r p l a n e f a l l s i n t h e Commercial Transpor t ca tegory.

S t e ~ 5: The fol lowing weight equa t ions apply t o t h e Ourania :

'struct : 1) Wing: Eqns (5 .6) and (5.7)

2 ) Adjustment f o r Fowler f l a p s : an ex t ra f a c t o r of 2 pe rcen t w i l l be added i n accordance wi th 5.2.2.2.

3 ) Empennage: Eqns (5.171, (5.181, (5.20)

4 ) Fuselage: Eqns (5.26) and (5.27)

5 ) Nacelles: Eqns (5 .35) and (5.37)

6) Landing Gear: Eqns (5.41) and (5.42)


2) Fuel system: Eqn. (6.24)

3 ) Propuls ion system: Eqns (6.241, (6.29)

4 ) Accessory d r i v e s , s t a r t i n g and i g n i t i o n system: Eqn. (6.34)

5 ) Thrus t reversers: Eqn. (6.36)

'feq: 1) F l i g h t c o n t r o l system: Eqns (7.51, (7.6)

Note: h y d r a u l i c s and pneumatics are i n- cluded i n item 1).

2) Electrical system: Eqns (7.151, (7.17)

3 ) Ins t rumenta t ion , a v i o n i c s and electro- n i c s : Eqns (7.23) and (7.25)

4 ) Air- condi t ion ing , pressurization and de- icing: Eqns (7.29) and (7.30)

5 ) Oxygen system: Eqns (7.35) and (7,371

6) APU: Eqn. (7.40)

7 ) Furn ish ings : Eqns (7.44) and (7.45)


8) Baggage and cargo handl ing: Eqn. (7.48) -

9 ) Opera t iona l items: See S e c t i o n 7.10. - =3

10) P a i n t : See S e c t i o n 7.15.

SteD The fol lowing l is t i temizes a l l requi red inpu t d a t a f o r e s t ima t ing t h e weight items l i s t e d i n s t e p s 2 and 5.

127,000 l b s nult = 2.5 S = 1,296 f t 2

'TO

VC = 295 k t s VD = 369 k t s nult = 3.75

A = 10 A = 0.32 A 1 /4 = 35 deg.

Sh = 254 f t ' bh = 35.6 f t t = 1.30 f t rh

A 114 = 45 deg.

- lf = 124.3 f t wf + hf = 26.4 f t qD = 461 psf

WL = 7,505 l b s nu l t e l = 4.0 df = 13.2 f t . P Z P 2 0 p s i l n P 1 1 . 7 f t = 28.3 f t

2 A i n l

Notes: 1) The va lue f o r nlim fo l lows from t h e V-n diagram of F igu re 4.5.

2 ) Most d a t a were obta ined from Ourania d a t a l i s t e d i n P a r t 11. The reader is reminded t h a t a d e t a i l e d geometr ic d e f i n i t i o n may be found i n P a r t I1 as Table 13.2, a Class I weight statement as Table 10.5. De ta i l ed d e f i n i t i o n s of l a y o u t s of f u s e l a g e , powerplant , wing, h igh l i f t system, empennage and landing gear may be found i n Chapters 4 ,5 ,6 ,7 ,8 and 9 r e s p e c t i v e l y .

S t e ~ 7: Tables 4.2a - 4 . 2 ~ l i s t a l l weights computed as p a r t of t h e Class I1 weight e s t i m a t i o n process .


S f e ~ 8: The Class I1 empty weight of t h e Ourania is 7 2 , 6 2 2 l b s . Th i s compares wi th 68,450 l b s f o r t h e Class I weight estimate. This r e p r e s e n t s a d i f f e r e n c e of 4 , 1 7 2 l b s which is 6 . 1 pe rcen t of t h e C l a s s I empty weight.

Seve ra l comments a r e i n o rde r :

1. an i t e r a t i o n through t h e equa t ions of S t e p 7 should be performed, t o determine t h e 'convergence' empty weight.

2. s e v e r a l weight s av ings can be made i n t h e Ourania:

a ) t h e empennage can be manufactured from composites which would y i e l d a weight saving of about 1 5 p e r c e n t , o r 359 l b s .

b ) t h e n a c e l l e s can be manufactured p a r t i a l l y from composites which would y i e l d a weight saving of about 1 0 p e r c e n t , o r 264 l b s .

c ) by manufacturing t h e low s t r e s s a r e a s of t h e wing and f u s e l a g e from composites, a weight sav ing of about 5 p e r c e n t should be f e a s i b l e . This would save 1 , 3 8 8 l b s .

d ) by us ing a quadruplex d i g i t a l f l i g h t c o n t r o l sys tem and us ing fly-by-wire i n s t e a d of mechanical f l i g h t c o n t r o l s . a weight sav ing of 1 5 p e r c e n t over t h e es t imated weight can be obta ined. T h i s would save 352 l b s .

e) by us ing l i t h i u m aluminum i n t h e primary wing and f u s e l a g e s t r u c t u r e , a weight sav ing of 6 pe rcen t is f e a s i b l e . This s aves 1,665 l b s .

By combining a) through e) a t o t a l weight sav ing of 4 , 0 2 8 l b s can be achieved. T h i s is c l o s e t o t h e d i sc repancy of 4 ,172 l b s . I t is t h e r e f o r e judged p o s s i b l e t o b r i n g t h e Ourania i n a t t h e o r i g i n a l l y es t imated empty weight of 68,450 l b s .

DS 9-10: Not needed, see i t e m e l , S t e p 8.

11: Th i s s t e p has been omi t ted t o save mace.


T a b l e 4 .2a C l a s s I1 Weight E s t i m a t e s f o r t h e O u r a n i a ....................................................

Component Methods : use as - =m

C l a s s I GD Torenbeek C l a s s I1 P a g e 11 E s t i m a t e

--------==--------------- ........................... -------- ---------------=---------------------------

S t r u c t u r e w e i g h t , Wst ruc t :

Wing 13 ,664 1 1 , 7 5 3 1 5 , 9 7 3 1 3 , 7 9 7

A d j u s t m e n t f o r Fowle r f l aps , 2 percent: 27 6

H o r i z . T a i 1 949 1 , 2 1 8 1 ,319 V e r t . T a i l 920 82 9 1 , 0 7 1 Empennage 3 ,253 1 , 8 6 9 2 ,047 2 , 3 9 0

F u s e l a g e 1 4 , 1 8 4 1 5 , 7 4 8 1 1 , 1 4 0 1 3 , 6 9 1

Nacelles 2,082 2 ,722 3 , 1 2 0 2 ,641

Nose Gear 573 7 83 716 Main Gear 4 ,632 4 , 2 0 8 3 ,904 Land ing g e a r 5 ,205 3 ,663 4 ,991 4 ,620 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Wstruc t 3 8 , 3 8 8 37 ,415 ---------------------------=------------------------- ........................... .........................

P o w e r p l a n t w e i g h t , W : -------------- ----------------- ----ewE------------=-----------------

E n g i n e s 9 ,224 9 , 2 2 4 9 ,224 9 , 2 2 4

F u e l system i n pwrplt 1 , 0 0 9 1 , 0 0 9

P r o p u l s i o n i n s t . 667 439 700

Acc .dr , S tar t , Ign 960

T h r u s t reversers 1 , 6 6 0 1 , 6 6 0 ------- ...................... --------=-------me---------------------

W 9 , 8 9 1 1 2 , 5 9 3 =gyg========----===============-------------==------- ------------ -------

P a r t V C h a p t e r 4 Page 56

Table 4.2b Class I1 Weight Estimates f o r t h e Ourania: PIPPIPI I I I~ I~PI~PIOJ~C.D==PI~P.~======~=~==~~=~=~===~======

Average Weight F r a c t i o n s f o r Fixed Equipment Breakdown ~ P P I I P I P I P ~ ~ P = ~ = P P P = = ~ X = X ~ X J = ~ : = = ~ ~ ~ P = P ~ ~ = ~ ~ = = = = = ~ P = I = ~

Note: t h e s e d a t a were used i n Table 4 . 2 ~ .

Component S i m i l a r Ai rp lane v p e : use as MCDD Boeing Class I1 DC-9-30 MD-80 737-200 727-100 E s t i m a t e

P = I I P e P t e ~ % = P = P I P I C = P o e = = = = = a ~ = r = = P P P P J e 5 = = ~ = = = = ~ = = = a = = = = =

Fixed equipment weight i t e m :

W ~ t t y p i c a l US a i r l i n e pa in t scheme: 0,0035

P P P P P P P P I I P O I I P I n P P P P P ~ e r r r = P = = % f : P ~ P e = ~ ~ l i = 6 I P = ~ = = = = = ~ = = = = =

i nc ludes h y d r a u l i c and pneumatic system

B t e : S p e c i f i c airplane t y p e d a t a from Tables A7.1a and A7.2a i n Appendix A.


Table 4 . 2 ~ Class I1 Weight Estimates for t h e Ourania _ I L _ ~ I / S J a _ a D = _ S = ~ ~ I ~ P L ~ ~ ~ 3 ~ ~ ~ P P = ~ ¶ a = = : = t ~ ~ = ~ s s ~ ~ ~ = 5 ~

use a s component - Methods: - Table 4.2b GD Torenbeek Class I1 ~127,000 Estimate - - - = - - - - _ _ D = _ _ _ = _ _ _ _ _ _ P S = _ _ I = I = a _ O J = 5 3 f ~ e . = s = 5 5 = ~ s s ----

Fixed equipment weight, Wfeq:

__pp_=D3__DL___P___==3:==D=-P'c3_=~I==5_sJ=_=_==~===~===s_=_~

Wfc 3,226 2,200 1,617 2,34 8

'bps : t h i s is included in Wfc

Wels 1,499 1,887 4,063 2,483

'iae 1,810 1,593 1,775 1,726

'api 1,737 4,251 2,166 2,718

Wax 191 241 210 2 14

W apu 8 81 1,016 1,016 9 82

'fur 8,928 7,467 7,565 7,9 87

Wbc 46 6 46 6 466

WoPs 3,245 3,245 3,245

W Table 4.2b: 445 .et..==s =====J====== s=-C5==I====II = J=_5== * s=======_s=s== W 21,517 19,121 22,123 22,614 .Srs..=.=.-= =-==s=-J P=P 5:11=55=== =====_===s=s===s

Class I1 empty weight, WE fo l lows from Eqn. (2.11 :

W~ = 37,415 + 12,593 + 22,614 = 72,622 l b s

Part V Chapter 4

S t e ~ 1: The fol lowing weight i tems a r e a l r eady known :

From Table 10 .6 , P a r t 11:

Payload weight: w~~ = 1 2 , 4 0 5 l b s

Crew weight: 'c r e w = 2 0 0 l b s

Fuel weight: 'F = 18,500 l b s

Trapped f u e l and o i l : W t f o = 300 l b s

From P a r t 11, p.140:

Engines, i n c l A/B: 'e = 6 , 0 0 0 l b s

& e ~ 2 : Weights need t o be es t imated f o r t h e fol lowing items:

1) Wing 2 ) Adjustment f o r Fowler f l a p s 3 ) Empennage

4 ) Fuselage 5 ) Tailbooms 6 ) Engine s e c t i o n

7 ) Landing Gear

t Weiat . Elpwr:

1) Engines 2 ) Af te rburners 3 ) Air i nduc t ion system

4 ) Fuel system 5 ) Propuls ion system

t w-t. Elfeg:

1) F l i g h t c o n t r o l s 2 ) E l e c t r i c a l system

3 ) Ins t rumenta t ion , a v i o n i c s and e l e c t r o n i c s

4 ) Ai r- condi t ion ing , p r e s s u r i z a t i o n and de- icing

5 ) Armament 6 ) Furn ish ings 7 ) Oxygen system

8 )Aux i l i a ry g e a r 9 ) GAU-8A Gun *

a e ~ 3: The s t r u c t u r a l arrangement drawing f o r t h e E r i s is p re sen ted i n Chapter 8 of P a r t 111.


S t e ~ 4: From a weight e s t ima t ing viewpoint t h i s a i r p l a n e f a l l s i n t h e F i g h t e r and Attack Ai rp lane category.

- =z

eD 5: The fol lowing weight e q u a t i o n s apply t o t h e E r i s :

Wst ruc t : 1) Wing: Eqn. (5.9)

2 ) Adjustment f o r Fowler f l a p s : an extra f a c t o r of 2 pe rcen t w i l l be added i n accordance wi th 5.2.2.2.

3 ) Empennage: Eqns ( 5 . 1 7 ) and (5.18)

4 ) Fuselage: Eqn. (5.26)

5 ) Tailbooms: Eqn. (5.27)

6) Engine s e c t i o n : See Class I, p. 1 4

7 ) Landing Gear: Eqns (5.41) and (5.42)


2 ) A i r i nduc t ion system: Eqn. (6 .9)

3 ) Fuel system: Eqn. (6.20)

4 ) P ropuls ion system: Eqns (6.231, (6.27)

Wfeq: The d a t a of Table 4.3b are used, i n addi- t i o n t o t h e fol lowing equat ions:

1) F l i g h t c o n t r o l system: Eqn. (7.11)

Note: h y d r a u l i c s and pneumatics a re inc luded i n i t e m 1).

2 ) Electrical system: Eqn. (7.19)

3 Ins t rumenta t ion , avionics and e l e c t r o- n i c s : Eqn. (7.25)

4 ) Air- condi t ion ing , p r e s s u r i z a t i o n and de- icing: Eqn. (7.33

5 ) Armament: Table 4.4b

6) Furn ish ings : Eqn. (7.47)


7 ) Oxygen sys tem: Eqn. (7.39)

8) A u x i l i a r y g e a r : T a b l e 4.4b

9 ) GAU-8A Gun: S e e P a r t I11 under weapons

-D 6: The f o l l o w i n g l is t i t emizes a l l r e q u i r e d i npu t d a t a f o r e s t i m a t i n g t h e we igh t items l i s t e d i n steps 2 and 5.

GW = 61,660 l b s nult = 11.0 S = 787 f t2

A 114 = 41 deg. *Th i s is f o r b o t h ver t ical t a i l s

I( inl = 1.25

lf = 41.3 f t hi = 6.83 f t f o r t h e f u s e l a g e

lf = 33.3 f t wf + hf = 3.06 f t f o r t h e booms

Sf gs = 2 ~ 3 0 . 6 = 61.2 f t2 fo r t h e booms

Nose gear: 1 2 0.06 0 0

Main g e a r : 33 0.04 0.021 0

N i n l = 2 Ld = 8 f t A i n l - 6.31 f t2

P2 = 30 ps i Kd = 1.0 Km = 1.0

Notes: 1) The value for nlim follows f rom the+-n d i a g r a m of F i g u r e 4.6.

2 ) Most d a t a were o b t a i n e d f rom E r i s d a t a l i s t e d i n P a r t 11. The r e a d e r is reminded t h a t a

P a r t V C h a p t e r 4 Page 6 1

d e t a i l e d geometr ic d e f i n i t i o n may be found i n P a r t I1 a s T a b l e 13 .3 , a C l a s s I weight s t a t emen t a s Table 10 .6 . ~ e t a i l e d d e f i n i t i o n s of l a y o u t s of f u s e l a g e , powerplant , wing,-high l i f t system, empennage and landing gear may be f o u n g i n Chapters 4 , 5 , 6 , 7 , 8 and 9 r e s p e c t i v e l y .

S f e ~ 7 : Tab le s 4.3a, 4.3b and 4 . 3 ~ l ist a l l weights computed a s p a r t o f t h e C la s s I1 weight e s t i m a t i o n process .

=D 8: The Class I1 empty weight of t h e E r i s is 3 5 , 7 5 5 l b s . T h i s compares wi th 33,500 l b s f o r t h e C l a s s I weight estimate. This r e p r e s e n t s a d i f f e r e n c e of 2 , 2 5 5 l b s which is 6.7 pe rcen t of t h e C l a s s I empty weight.

Seve ra l comments a r e i n o rde r :

1. an i t e r a t i o n through t h e e q u a t i o n s of S t ep 7 should be performed, t o determine t h e 'convergence' empty weight.

2. s e v e r a l weight s av ings can be made i n t h e E r i s :

a ) t h e e n t i r e primary s t r u c t u r e can be made from composites. Th i s could y i e l d a p o t e n t i a l s av ings of 1 0 p e r c e n t o r 2 , 2 4 6 l b s .

b ) by us ing a quadruplex d i g i t a l f l i g h t c o n t r o l system and us ing fly-by-wire i n s t e a d of mechanical f l i g h t c o n t r o l s , a weight saving of 1 5 p e r c e n t over t h e estima- t e d weight can be obta ined. Th i s would s a v e 2 5 4 l b s .

By combining a ) and b ) a t o t a l weight saving of 2 , 5 0 0 l b s can be achieved. I t is t h e r e f o r e judged p o s s i b l e t o b r i n g t h e E r i s i n a t a weight below t h e o r i g i n a l l y es t imated empty weight.

~ t e ~ s 9- 10: Not needed, see i t e m e l , S t e p 8.

e~ 11: Th i s s t e p has been omi t ted t o save space.


Table 4 . 3 a Class I1 Weight Estimates f o r t h e E r i s = I = = I = P P P = = = P = P P I = P = = 5 : = P t = = P = = a = = = = = = = = = = ~ = = ~ = = = =

Component Methods : Use as C las s I1 GD Torenbeek C l a s s I1 Page 1 4 E s t i m a t e

=='==X==PIIPtPPP=PI1==PP=O='=I=====i===================~a===u===

S t r u c t u r e weight, Wstruct : =I==PPD=P=======PPSCP==~=======a========a============

Wing 6 , 7 6 2 9 , 4 9 0 8 ,126

Adjustment f o r Fowler f l a p s , 2 p e r c e n t : 1 6 3

Horiz.Tai1 7 2 0 Vert .Tai1 9 3 8 Empennage 1 , 5 9 7 1 , 6 5 8

Fuselage 7 ,347 5 , 0 4 4 i n c l . booms

Booms 4 5 8

Engine Sec t ion 1 6 0 1 6 0

Nose Gear 554 2 6 7 443 Main Gear 2 , 2 1 4 1 , 6 0 3 1 , 7 6 8 Landing g e a r 2 , 7 6 8 1 , 9 9 6 1 , 8 7 0 2 , 2 1 1 ===='=======DPP==tt==='IDP=SPPPC=rP5==Pr==========a==

W s t r u c t 1 8 , 6 3 4 2 0 , 3 0 4 1 8 , 7 1 3 P P = = t = Z = P = = = P P P P S = = = = = P - - - - - - - - P P = = L a P I = = = = =

Powerplant weight, W : I=~===P===P=I I I IP=P=ex~==~L:= iC~~~=~========~=====~=~=P Engines 4 , 0 0 0 4 , 0 0 0 4 , 0 0 0 4 , 0 0 0

Af te rbu rne r s 2 , 0 0 0 2 , 0 0 0 2 , 0 0 0 2 , 0 0 0

Air ind. s y s t . i n propuls . 4 4 5 4 4 5

Fuel system i n propuls . 7 7 7 7 7 7

Propuls ion i n s t . 2 , 8 3 4 * * 7 8 8 4 5 * * * a=PI==P'====IS==sIP=========5=P==LIL:==a5==============

W 8, 834 7 , 6 3 2 8 ,067 3BYg===a=======a=t===5==5============0=============== * 1 / 2 ( 7 , 3 4 7 - 4 5 8 + 5 , 0 4 4 ) = 5 9 6 7

3?

**inc ludes a i r i nduc t ion and f u e l system


T a b l e 4.3b Weight F r a c t i o n E s t i m a t e s f o r t h e E r i s : ===P====PP=tP==============~===============~======

A v e r a g e Weight F r a c t i o n s f o r F i x e d Equipment Breakdown -

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

-

Note : ' these d a t a were u s e d i n T a b l e 4 . 3 ~ .

Component S i m i l a r Airplane Type: U s e as R e p u b l i c Chance Vought Grumman C l a s s I1 F105B F 8U A2F(A6) E s t i m a t e

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

F i x e d e q u i p m e n t w e i g h t item:

Wf"r 0.0077 0.0069 0.0137 0.0094

'aux 0.0029 0.0060 0.0045 =XI===~==o===PPPI==1PIIBS=PIIP===XIP==================================

i n c l u d e s h y d r a u l i c and pneumatic sys tem

Note: Specific airplane type d a t a f r o m T a b l e s A9.2a, A9.3a a n d A9.4a i n Appendix A.


Table 4 . 3 ~ Class I1 Weight Estimates f o r t h e Eris PPPISI=DPPPS==PPISP=IIP=P==P~==~P~=P=~:==~=~===%====

Component Methods : U s e as Tab le 4.3b GD Torenbeek C l a s s I1 ~ 6 1 , 6 6 0 E s t i m a t e

----==- ---- -PIPPIPP==~=PPS=PPI===CIPP~PP====~==~===OP==========

Fixed equipment weight , Wfeq:

- Whps: t h i s is inc luded i n Wfc

GAU-8A Gun 2,014 P a r t 11, Table 10.6: 2,014 S P P I I I P I P P P = I I P I P I = = = = = P = = = P ~ = t = e = - i : = = = 3 C = = = u = = = = I = = = = ~ n ~

T h i s d i s a g r e e s s i g n i f i c a n t l y w i t h Wfeq i n Tab le 10.6 of P a r t 11.

Class I1 empty weight , WE f o l l o w s from Eqn.(2.1):

WE = 18,713 + 8,067 + 8,975 = 35,755 l b s


5. CLASS I1 METHOD FOR ESTIMATING STRUCTURE WEIGHT ====IS====X=I=P=5==PI=PI===~===X=z=t=t=======z=========

The a i r p l a n e s t ruc ture weight, Wstruct w i l l . be assu-

med t o cons i s t of t h e fol lowing components:

5.1 Wing, Ww

5.3 Fuselage, Wf

5.2 Empennage, W emP

5.4 Nacelles, W,

5.5 Landing g e a r , W 4

Therefore:

Wstruct ww + W emp

+ W f + w n + W 9

(5.1)

Equations f o r s t ruc ture weight est imation are p re sen ted f o r t h e fol lowing t y p e s of airplanes:

1. General Avia t ion Ai rp lanes 2. Commercial Transpor t A i rp l anes 3. M i l i t a r y P a t r o l , Bomb and Transpor t Ai rp lanes 4. F i g h t e r and A t t a c k A i rp l anes

The fol lowing equat ions should be app l i ed on ly t o small, r e l a t i v e l y low performance t y p e a i r p l a n e s wi th maximum speeds below 200 k t s . The e q u a t i o n s apply t o wings of two types :

Cant i lever wings: Eqn. (5.2) S t r u t braced wings: Eqn. (5.3

~ o t h equat ions include: weight of wing t i p f a i r i n g wing control s u r f a c e s

~ o t h equations exclude: f u e l t a n k s wing I f u se l age spar ca r ry - through structure e f f e c t of sweep angle

For canti lever winas: 0.397 ( s ) O . 360 0.397 712 (5 .2) W, = 0.04674(WT0) ("u l t )

>


-

w ~ O = take-of f weight i n l b s , 2 S = wing a r e a i n f t ,

"u l t = design u l t ima te load f a c t o r

A = wing aspect r a t i o

Note t h a t Eqn. (5 .3 ) does not account f o r WTO. It

should t h e r e f o r e be used with caut ion. The reader should a l s o r e a l i z e t h a t wings i n t h i s category have maximum th ickness r a t i o s of around 1 8 percent .

5.1.1.2 USAF Method

The following equat ion a p p l i e s t o l i g h t and u t i l i t y type a i r p l a n e s with performance up t o about 300 k t s :

of new terms:

A 114 = wing q u a r t e r chord sweep angle

A = wing t a p e r r a t i o

(tic), = maximum wing th ickness r a t i o

V, = maximum l e v e l speed a t s e a l e v e l i n k t s

5.1.1.3 T o r m e e k Method

The following equat ion a p p l i e s t o l i g h t t r a n s p o r t a i r p l a n e s w i t h take-off weights below 1 2 , 5 0 0 l b s :

See s p e c i a l notes i n Sect ion 5.2.2.

part v Chapter 5 Page 6 8

of new t e rms :

b = wing span i n f t

I 1 1 2 = wing semi-chord sweep ang le

tr = maximum t h i c k n e s s of wing r o o t chord i n f t

Note: This equa t ion is v a l i d on ly i n t h e fol lowing parameter ranges:

M B from 0.4 t o 0.8, ( t l c ) , from 0 .08 t o 0.15,

and A from 4 t o 12.

of new term:

M H = maximum Mach number a t s e a l e v e l

The fol lowing equa t ion a p p l i e s t o t r a n s p o r t a i r p l a n e s w i th take- off weights above 12,500 l b s :

W~~~ = maximum ze ro f u e l weight = WTO - WF (5 .8)

1. Eqns. (5.6) and (5.7 i n c l u d e t h e weight o f normal high l i f t dev ices as w e l l as a i l e r o n s .

2. For s p o i l e r s and speed b rakes 2 p e r c e n t should be added.

p a r t v Chapter 5 Page 69

I f t h e a i r p l a n e has 2 wing mounted engines reduce t h e wing weight by 5 percen t . I f t h e a i r p l a n e has 4 wing mounted engines reduce t h e wing weight by 10 percen t . I f t h e land ing g e a r is no t mounted under t h e wing reduce t h e wing weight by 5 percen t . For braced wings reduce t h e wing weight by 30 percen t . The r e s u l t i n g wing weight estimate does i n c l u d e t h e weight of t h e s t r u t . The l a t t e r is roughly 10 pe rcen t of t h e wing weight. For Fowler f l a p s add 2 p e r c e n t t o wing weight.

For p r e d i c t i n g wing weight it is suggested t o u se Eqns. (5.6) and (5.7) bu t wi th t h e a p p r o p r i a t e va lue f o r " u l t ' For t h i s t y p e of m i l i t a r y a i r p l a n e t h e u sua l va lue

fo r n u l t is 4.5. Refer t o Table 4 .1 f o r a l i s t i n g of m i -

l i t a r y l i m i t l oad f a c t o r s .

Note: wing weight i n m i l i t a r y a i r p l a n e s is o f t e n based on t h e f l i g h t des ign g r o s s weight , GW, r a t h e r t h a n w ~ o ' Check t h e miss ion s p e c i f i c a t i o n and/or t h e app l i ca-

b l e m i l i t a r y s p e c i f i c a t i o n s t o de te rmine which weight va lue t o use i n Eqns. ( 5 . 6 ) and ( 5 . 7 ) .

5 .1 .4 .1 GD Method

t e r and attack a i w e s :

er and attack a


of new te rms:

= 1.00 f o r f i x e d wing a i r p l a n e s and 'W = 1 . 1 7 5 f o r v a r i a b l e sweep wing a i r p l a n e s

A ,, = l e ad ing edge sweep ang le of t h e wing

Note: wing weight i n m i l i t a r y a i r p l a n e s is o f t e n based on t h e f l i g h t des ign g r o s s weight , GW, r a t h e r t han WTo. Check t h e miss ion s p e c i f i c a t i o n andlor t h e app l i ca-

b l e m i l i t a r y s p e c i f i c a t i o n s t o de te rmine which weight t o use i n Eqns. ( 5 . 9 ) and (5 .10) .

Empennage weight, Wemp w i l l be expressed a s fol lows:

we^ P Wh + WV + WCI ( 5 . 11 )

where: Wh - h o r i z o n t a l t a i l weight i n l b s

WV = v e r t i c a l t a i l weight i n l b s

Wc - canard weight i n l b s

Equations f o r empennage weight components a r e p re sen ted i n t h e remainder of t h i s s e c t i o n .

The fol lowing equa t ions should be app l i ed on ly t o small, r e l a t i v e l y low performance t y p e a i r p l a n e s wi th maximum speeds below 200 k t s .

3 .184 (WTO) 0. 887(Sh)0e 10 l (Ah) 0.138 -

Wh - ................................. 0.223 (5 .12)

5 7 . 5 ( t rh

Note t h a t no f a c t o r f o r h o r i z o n t a l t a i l sweep is included.

Yerti-1 tail: *


Eanard: For a l i g h t l y loaded canard, Eqn. (5 .12) may be used. For a s i g n i f i c a n t l y loaded canard (such a s on t h e ~ P 1 8 0 and t h e S t a r s h i p I) it is suggested t o use t h e appropr ia te wing weight equation.

of t e r m

w ~ o = take-off weight i b l b s

Sh = hor izon ta l t a i l a r e a i n f t 2

Ah = hor izon ta l t a i l aspec t r a t i o

t = hor izon ta l t a i l maximum r o o t th ickness i n f t rh

Sv = v e r t i c a l t a i l a r e a i n f t 2

A, = v e r t i c a l t a i l aspec t r a t i o

t = v e r t i c a l t a i l maximum r o o t th ickness i n f t r~

l / 4 v = v e r t i c a l t a i l q u a r t e r chord sweep angle

5.2.1.2 U w Method


o n t a l tail:

-.

Note t h a t sweep angle is no t a f a c t o r i n t h i s equation.

V e r t i c a l tail:

Wv = 98. 5( (WTonUlt / l~5)o~ 8 7 ( ~ V / 1 0 0 ) 1 * 2 x

x0.289(bv/t ) 0.510.4S8 (5 .15) rv

Again, sweep angle is no t a f a c t o r i n t h i s equation.

The comments made under 5.2.1.2 a l s o apply.


of new terros:

lh = d i s t a n c e from wing c l 4 t o hor. t a i l Gh/4 i n f t

bh = hor izon ta l t a i l span i n f t

bv = v e r t i c a l t a i l span i n f t

The following equat ion app l i e s t o l i g h t t r a n s p o r t a i r p l a n e s w i t h design d i v e speeds up t o 250 k t s and w i t h conventional t a i l conf igura t ions :

I f t h e a i r p l a n e a l s o has a canard, t h e comments made under 'canard' i n 5 .2 .1 .2 a l s o apply here.

Note: sweep angle is n o t a f a c t o r i n t h i s equation.

V e r t i c a l t a k

Canard: Comments made under 5.2.2.2 a l s o apply here.

of new t-

zh = d i s t a n c e from t h e v e r t i c a l t a i l roo t t o where t h e hor izon ta l t a i l is mounted on t h e v e r t i c a l t a i l , i n f t . f o r fuse lage munted hor izon ta l t a i l s , s e t zh = 0.


-1, = d i s t . from wing GI4 t o v e r t . t a i l cv14 i n it

= rudder a r e a i n it 2

A V = v e r t i c a l t a i l t a p e r r a t i o

The following equat ion a p p l i e s t o t r a n s p o r t a i r p l a n e s and t o business jets w i t h design d i v e speeds above 250 k t s .

where K h t a k e s on t h e following values:

K h = 1.0 f o r f ixed incidence s t a b i l i z e r s

Kh = 1.1 f o r v a r i a b l e incidence s t a b i l i z e r s

Y e r t i d t U

wv =u (5 .20)

= KvSv[3. 81 ~ ( S ~ ) ~ ~ ~ V ~ / ~ . ~ O ~ ( C O ~ l12v ) ' I2) - 0.2871

where Kv t akes on t h e following values:

Kv = 1.0 f o r fuse lage mounted hor izon ta l t a i l s

f o r f i n mounted hor izon ta l t a i l s :

of new terms:

VD = design dive speed i n KEAS

1/Zh hor izon ta l t a i l semi-chord sweep angle

l l Z V v e r t i c a l t a i l semi-chord sweep angle

Canardt The comments made under 5.2.2.2 a l s o apply here.


5.2.3 P a t r o l , T r v

See Sub-section 5.2.4.

For es t imat ion of empennage weight of a i r p l a n e s i n t h i s category, use t h e methods of sub- section 5.2.2. Be s u r e t o use t h e proper va lues f o r u l t i m a t e load f a c t o r . See Table 4.1.

Note: empennage weights of m i l i t a r y a i r p l a n e s a r e o f t e n based on t h e f l i g h t design g r o s s weight, GW, r a t h e r than WTO. Check t h e mission s p e c i f i c a t i o n andlor t h e

app l i cab le m i l i t a r y s p e c i f i c a t i o n s t o determine which weight t o use.

The equat ions presented f o r fuse lage weight es t imat ion a r e v a l i d f o r land-based a i r p l a n e s only. For f l y i n g boats and amphibious a i r p l a n e s it is suggested t o mult iply t h e fuse lage weight by 1.65:

1.65Wf W f f l . boat

For f l o a t equipped a i r p l a n e s t h e weight due t o t h e f l o a t s may be found with Eqn.(5.27), by s u b s t i t u t i n g f l o a t wetted a r e a f o r Sfgs.

For es t imat ion of tailboom weight it is suggested t o use Eqn.(5.27) appl ied t o each tailboom ind iv idua l ly , but with Kf = 1.

5 .3 .1 G w Aviation . .

The following equat ions should be appl ied only t o smal l , r e l a t i v e l y low performance type a i r p l a n e s with maximum speeds below 200 k t s .


WTO = take-of f weight i n l b s

N Pax

= number of passengers

= fuse lage length , no t including nose mounted If-' n a c e l l e length i n f t

m t e s : 1. These equat ions do n o t account f o r pressur ized fuselages.

2. There is no explanat ion f o r why t h e fu- s e l a g e weight of low wing a i r p l a n e s does n o t depend on t h e number of passengers.

3. For t h i s type a i r p l a n e t h e crew is coun- t e d i n t h e number of passengers.


of new terlrrs:

" u l t = u l t i m a t e load f a c t o r

lf = fuse lage length i n f t

wf = maximum fuse lage width i n f t

hf = maximum fuse lage he ight i n f t

VC = design c r u i s e speed i n KEAS

p a r t v Chapter 5 Page 7 6

The f a c t o r Kinl t akes on t h e following values:

= 1 . 2 5 f o r a i r p l a n e s with i n l e t s i n o r on t h e fuse lage f o r a buried engine i n s t a l l a t i o n

K i n l = 1.0 f o r i n l e t s loca ted elsewhere

of new term: - qD = design d ive dynamic p ressu re i n psf

The following equat ion a p p l i e s t o t r a n s p o r t a i r p l a n e s and t o business j e t s with design d ive speeds above 250 k t s .

112 Wf = Om021Kf((VDlh/ ( ~ f + h f ) 1 (SfgS ) 1.2 ( 5 . 2 7 )

The cons tant Kf t a k e s on t h e following values:

Kf = 1.08 f o r a pressur ized fuse lage

= 1 .07 f o r a main gear a t tached t o t h e fuselage.

= 1.10 f o r a cargo a i r p l a n e with a cargo f l o o r

These e f f e c t s a r e m u l t i p l i c a t i v e f o r a i r p l a n e s equipped with a l l of t h e above.

of new terms:

VD = design d ive speed i n KEAS

lh = d i s t a n c e from wing E l 4 t o hor. t a i l Eh/4 i n f t

= fuse lage g ross s h e l l a r e a i n f t 2 S fgs

For USAF a i r p l a n e s , Eqn. (5 .26 ) may be used.

For USN a i r p l a n e s t h e following equat ion should a be used :


Values f o r Kinl a r e as g iven i n 5.3.2.1.

n of new t e u - - qL = des ign d i v e dynamic p r e s s u r e i n p s i

For e s t i m a t i o n of f u s e l a g e weights Equat ions (5 .26 ) o r ( 5 . 2 8 ) may be used.

-a: I n us ing Eqn. (5 .26 ) f o r f i g h t e r s , l e a v e nff t h e f a c t o r 2 a t t h e beginning of t h e equat ion.

The n a c e l l e weight is assumed t o c o n s i s t of t h e fol lowing components :

1. For podded engines: t h e s t r u c t u r a l weight a s s o c i a t e d wi th t h e engine e x t e r n a l d u c t s and o r cowls. Any pylon weight is included.

2. For p r o p e l l e r d r i v e n a i r p l a n e s : t h e s t r u c t u r a l weight a s s o c i a t e d wi th t h e engine e x t e r n a l d u c t s and o r cowls p l u s t h e weight due t o t h e eng ine mounting t r u s s e s .

3. For bu r i ed engines: t h e s t r u c t u r a l weight a s s o c i a t e d wi th s p e c i a l cowling and o r duc t ing p r o v i s i o n s ( o t h e r t h a n t h e i n l e t duc t which is included i n t h e a i r i nduc t ion system under powerplant weight , S e c t i o n 6 .2 ) and any s p e c i a l eng ine mounting p rov i s ions .

5 .4 .1 General Aviat- . .

The fol lowing equa t ions should be app l i ed on ly t o s m a l l , r e l a t i v e l y low performance t y p e a i r p l a n e s wi th maximum speeds below 200 k t s .

wn = K n w ~ ~ ( 5 . 2 9 )

The c o n s t a n t Kn takes on t h e fo l lowing va lues :

Kn = 0.37 l b s l h p f o r r a d i a l eng ines

Kn = 0.24 l b s l h p f o r h o r i z o n t a l l y opposed engines


WTO = take-of f weight i n l b s

These d a t a should n o t be app l i ed t o t u r b o p r o p e l l e r n a c e l l e s .

I n t h i s method, t h e n a c e l l e weight is included i n t h e powerplant weight: r e f e r t o Chapter 6.

For s i n a l e e r>ror>ellerdriven ai-es wikb t h e f w a e nose:

This weight i n c l u d e s t h e e n t i r e engine s e c t i o n forward of t h e f i r e w a l l .

F U t i - l t i e w i t h t o n en-

Wn = 0.32PT0 f o r h o r i z o n t a l l y opposed engines (5 .31)

Wn = 0.045(PT0) 'I4 f o r r a d i a l eng ines (5 .32 )

Wn = 0.14(PT0) f o r tu rboprop engines (5 .33)

w e s : 1. S ince PTo is t h e t o t a l requ i red take- off

horsepower, t h e s e weight estimates inc lude t h e weights of & n a c e l l e s .

2. I f t h e main land ing g e a r r e t r a c t s i n t o t h e n a c e l l e s , add 0.04 l b s f h p t o t h e n a c e l l e weight

3. I f t h e engine exhaus t s over t h e wing, as i n t h e Lockheed Electra, add 0.11 l b s f h p t o t h e n a c e l l e weight.

o i e t en-

P a r t V Chapter 5

-t

( 5 . 34 )

Page 79

- n of terms:

*in1 = number of i n l e t s

= cap tu re a r e a per i n l e t i n f t 2 *in1

1, = n a c e l l e length from i n l e t l i p t o compressor face in f t

P2 = maximum s t a t i c p ressu re a t engine compressor f a c e in p s i . Typical va lues range from 1 5 t o SO p s i .

For t u r b o l e t or low b v ~ a s s r a t i o turbofan e n d r u % ~

Since TTO is t h e t o t a l required take-off t h r u s t .

t h e s e equat ions account f o r t h e weight of U nace l l e s .

5.4.3 Mi-v P a t r o l B o m b T w ~ o r t A i w

For a l l a i r p l a n e s in t h i s category Eqns. (5 .34) and (5 .35) may be used.

5 .4 .4 Fiqhfer and Bftack

For a l l a i r p l a n e s i n t h i s category Eqns. (5 .34 ) and (5 .35) may be used.

The. following equat ions should be appl ied only t o small , r e l a t i v e l y low performance type a i r p l a n e s with maximum speeds below 2 0 0 k t s .


wheels + t i r e s m.g. s t r u t assemb1ym.g.

wheels + t i res n.g. s t r u t assembly n.g.

'TO = take- of f weight i n l b s

WL = d e s i g n l and ing weight i n l b s (See Tab le 3.3. P a r t I f o r d a t a r e l a t i n g WL t o wTO)

= u l t i m a t e load f a c t o r f o r l and ing . may be t aken as 5.7

1, = shock s t r u t l e n g t h f o r main g e a r i n f t m

1 = shock s t r u t l e n g t h f o r nose g e a r i n f t 'n

$.5.1.2 W Method

The fo l l owing e q u a t i o n a p p l i e s t o l i g h t and u t i l i t y t y p e a i r p l a n e s w i t h performance up t o about 300 k t s :

0. ) O . 684 w!3

= 0 . 0 5 4 ( 1 ) 'm L u l t . 1

Notes: 1) T h i s e q u a t i o n i n c l u d e s nose g e a r weight .

2 , N u l t . l may be t a k e n a s 5.7.

5.5.2.2 T m e e k Method

The fo l lowing e q u a t i o n a p p l i e s t o t r a n s p o r t a i r p l a n e s and t o b u s i n e s s jets w i t h t h e main g e a r mounted


on t h e wing and t h e nose g e a r mounted on t h e f u s e l a g e : -

= (Ag + B ( W 3 /4 + CgWTO + Dg (WTO) 3 / 2 ) Wg Rgr g TO

(5.42)

?he f a c t o r K t a k e s on t h e fo l l owing va lues : g r

K = 1.0 f o r low wing a i r p l a n e s gr

K = 1.08 f o r h igh wing airplanes 4,

The cons tan t s A through D a r e d e f i n e d i n g 53

Tab le 5.1 which is t a k e n from Refe rence 14.

Tab le 5.1 Constants i n Landing Gear Weight Eqn.(5.42) --------=--- -------- -^-PP3PIP==PIa====t=====3:P=====5=3==a======~

A i r p l a n e Gear Gear A 9 Bg Cg

D TYPe Type Comp.

4

Je t T r a i n e r s R e t r . Main 33.0 0.04 0.021 0.0 and B u s i n e s s Nose 12.0 0.06 0.0 0.0 J e t s

Other c i v i l Fixed Main 20.0 0.10 0.019 0.0 airplanes Nose 25.0 0.0 0.0024 0.0

T a i l 9 0.0 0.0024 0.0 Retr. Main 40.0 0 . 1 0.019 1.5~101:

Nose 20.0 0.10 0.0 2 . 0 ~ 1 0 T a i l 5.0 0.0 0.0031 0.0

For USAF airplanes, Eqns. (5.41) and (5.42) may be used.

For USN airplanes t h e fo l l owing equation should be used:

For USAF airplanes, Eqns. (5 .41) and (5.42) may be used.

For USN a i r p l a n e s , Eqn.(5.43) shou ld be used.


6. CLASS I1 METHOD FOR ESTIMATING POWERPLANT WEIGHT ==PP=====PI=s=P==P===X====X========P===============

The a i r p l a n e powerplant weight, W Pwr

w i l l be assumed

t o c o n s i s t of t h e fol lowing components:

6.1 Engines, We: t h i s i nc ludes engine, exhaus t ,

coo l ing , supercharger and l u b r i c a t i o n systems.

Bote: a f t e r b u r n e r s and t h r u s t r e v e r s e r s a r e n o t always inc luded under engines . They a r e o f t e n t r e a t e d as a s e p a r a t e powerplant component.

6.2 Air i nduc t ion system, W : t h i s i nc ludes i n l e t

d u c t s o t h e r t han n a c e l l e s , ramps, s p i k e s and a s s o c i a t e d c o n t r o l s .

6.3 P r o p e l l e r s , W Prop

6 . 4 Fuel System, W f s

6 . 5 Propuls ion System, W t h i s i nc ludes : P '

*engine c o n t r o l s * s t a r t i n g systems * p r o p e l l e r c o n t r o l s *p rov i s ions f o r engine i n s t a l l a t i o n

Note: i n s t e a d of t h e words 'p ropuls ion sys tem' , t h e words 'p ropuls ion i n s t a l l a t i o n ' o r even 'engine i n s t a l l a t i o n ' a r e sometimes used.

Therefore :

a 1 Note: f o r powerplant weight p r e d i c t i o n s it is h igh ly recommended t o o b t a i n a c t u a l weight d a t a from engine manufacturers.

Equat ions f o r powerplant weight p r e d i c t i o n a r e p resen ted f o r t h e fol lowing t y p e s of a i r p l a n e s :

1. General Avia t ion Ai rp lanes 2. Commercial Transpor t A i rp l anes 3. M i l i t a r y P a t r o l , Bomb and Transpor t Ai rp lanes 4. F i g h t e r and Attack Ai rp l anes


The following equat ions should be appl ied only t o small , r e l a t i v e l y low performance type a i r p l a n e s with maximum speeds below 200 k t s .

= K p P ~ ~ ( 6 . 2 )

The f a c t o r K t akes on t h e following values: P

For p i s t o n engines: K = 1.1 t o 1.8, depending on P

whether o r no t supercharging is used.

For turbopropel le r engines: K = 0.35 t o 0.55. P

These weights represent t h e so- cal led engine dry weight. Normal engine accessor i e s a r e included i n t h i s weight but engine o i l is not.

We = weight of a l l engines i n l b s

'TO = required take-off power i n hp

6.1.1.2 U W Method

+ W + W = 2.575(W ) 0. 92zN

+ 'ai P e ng e ( 6 . 3 ) Prop

U s e engine manufacturers d a t a t o obta in W o r use Eqn. ( 6 . 2 ) . eng

of new t . e r m

W eng

= weight per engine i n l b s

Ne = number of engines

E Q r ~ r o ~ e l l e r dr iven W e s :

WpWK - KW (We + 0 . 24PTO) ( 6 . 4 )

The cons tant K t akes on t h e following values: Pg


K = 1 . 1 6 f o r s i n g l e engine t r a c t o r i n s t a l l a t i o n s Pg

K ~ g 1 . 3 5 f o r mult i- engine i n s t a l l a t i o n s

For superchargers t h e fol lowing a d d i t i o n a l weight is incu r r ed :

wpwr = KpgKthrWe ( 6 . 6 )

The c o n s t a n t K t a k e s on t h e fol lowing va lues : Pg

K = 1 . 4 0 f o r a i r p l a n e s wi th bur ied engines W

The c o n s t a n t Kth r takes on t h e fol lowing va lues :

K t h r = 1.00 f o r a i r p l a n e s wi thout t h r u s t r e v e r s e r s

K t h r = 1.18 f o r a i r p l a n e s wi th t h r u s t r e v e r s e r s

6.1.2 C o m T r q a ~ ~ o r t A i w

U s e of a c t u a l eng ine manufactures d a t a is h igh ly recommended. F igu re 6 . 1 p rov ides a g r a p h i c a l summary of engine dry weights v e r s u s take- off t h r u s t . F igu re 6.2 g i v e s a g r a p h i c a l summary of engine d ry weights ve r sus take- off s h a f t horsepower.

When using F igu res ( 6 . 1 ) o r (6.21, keep i n mind t h a t :

- 'e - NeWeng ( 6 . 7 )

where W eng is t h e weight p e r engine.

Equat ions ( 6 . 5 ) and ( 6 . 6 ) may a l s o be used t o o b t a i n an i n i t i a l estimate.

See Sub- section 6.1.2.

See Sub- section 6.1.2.


Fiaure 6 . 1 Turbo-: Take - o f f T b u s t a u W W e i a h t Trends

F i a u e 6 . 2 Tu-roos: me - a f f t Horse Power Drv weight Trends

Part V Chapter 6 Page 86

6.2.1 General Avia t ion AirPlanes . .

6.2.1.1 C-

Wai is included i n t h e p ropu l s ion system weight. W P' 6.2.1.2 Methpd

See 6.2.1.1.

6.2.1.3 T o m e e k Me-

The a i r i nduc t ion system weight is s p l i t i n t o two items: t h e f i r s t one f o r d u c t suppor t s t r u c t u r e . t h e second one f o r t h e subsonic duc t l ead ing from t h e i n l e t l i p t o t h e engine compressor face .

(duc t suppor t s t r u c t u r e )

( subsonic p a r t of d u c t )

The f a c t o r s Kd and K, are de f ined as fo l lows:

Kd - 1.33 f o r d u c t s wi th f l a t c r o s s s e c t i o n s

1.0 f o r d u c t s wi th curved c r o s s s e c t i o n s

Km - 1.0 f o r MD below 1.4

= 1.5 f o r MD above 1.4

Ld - duc t l e n g t h i n f t

N i n l = number of i n l e t s


= c a p t u r e a r e a p e r i n l e t i n f t 2 A i n l - P2 = maximum s t a t i c p r e s s u r e a t engine compressor - - f a c e i n p s i . Typica l v a l u e s range from 1 5 t o

50 p s i .

For podded enqJne i n s t a l l a t i o n s :

The a i r i nduc t ion system weight is inc luded i n t h e n a c e l l e weight, Wn.

The c o n s t a n t Kd takes on t h e fol lowing va lues :

Kd = 1.0 f o r d u c t s wi th curved c r o s s s e c t i o n s

1 .33 f o r d u c t s w i th f l a t c r o s s s e c t i o n s

The a i r i nduc t ion system weight is included i n t h e n a c e l l e weight, Wn.

W t e : For supe r son ic i n s t a l l a t i o n s a d d i t i o n a l weight items due t o t h e s p e c i a l i n l e t requirements a r e needed. See Sub- section 6.2.4.

6.2.3. u v - PatrobBomb and Tr-t

See S e c t i o n 6.2.2.

6 .2 .4 .1 GD Method

For p r e d i c t i o n of t h e duc t suppor t s t r u c t u r e weight and t h e d u c t weight, Eqn. (6 .9 may be used.

P a r t i c u l a r l y i n supe r son ic a p p l i c a t i o n s t h e fol lowing a d d i t i o n a l weight items due t o i n l e t p r o v i s i o n s may be incur red :


The f a c t o r K r t akes on t h e following values:

K r = 1.0 f o r MD below 3.0

= (MD + 2115 f o r MD above 3.0

of new term:

Lr is t h e ramp length forward of t h e i n l e t t h r o a t i n f t

Wsp = Ks (Ninl ) (Ainl 1 ( 6 . 1 2 )

The cons tant Ks t a k e s on t h e following values:

Ks = 12 .53 f o r h a l f round f ixed sp ikes = 15 .65 f o r f u l l round t r a n s l a t i n g sp ikes = 51. 80 f o r t r a n s l a t i n g and expanding sp ikes

Note: t h e s e weights a l s o apply t o supersonic commercial i n s t a l l a t i o n s .

6 .3 .1 General A v w n AirDlanes . .

I t is recommended t o use p r o p e l l e r manufacturer d a t a whereever poss ib le . Lacking a c t u a l d a t a t h e equat ion of Sub-section 6.3.2 may be used.

Appendix A con ta ins p rope l l e r i n s t a l l a t i o n d a t a f o r a number of a i rp lanes . P rope l l e r i n s t a l l a t i o n weights usua l ly inc lude t h e p r o p e l l e r con t ro l s .

The cons tant K prop1

takes on t h e following values: 3

K prop1

= 24.0 f o r turboprops above 1 , 5 0 0 shp

= 31.92 f o r p i s t o n engines and f o r turboprops below 1 ,500 shp


N is t h e number of p r o p e l l e r s - P Nbl is t h e number of blades pe r p r o p e l l e r

D is t h e p rope l l e r diameter i n f t P

PTO is t h e required take-off power i n hp

Ne is t h e number of engines

The f a c t o r K prop2

t a k e s on t h e following values:

K prop2

= 0.108 f o r turboprops

K prop2

= 0 . 1 4 4 f o r p i s t o n engines

The reader is asked t o show t h a t equat ions (6 .13) and ( 6 . 1 4 ) a r e i n f a c t t h e same.

6.3.3 M i l i t a r y Patrol., Bomb and Traasr>ort A i mlam&



6.4 FUEL -EM W E I W ESTIMATION

m t e : I n some a i r p l a n e s t h e f u e l system is used t o con t ro l t h e cen te r of g r a v i t y loca t ion . Airplanes with relaxed s t a t i c s t a b i l i t y andlor supersonic c r u i s e a i r p l a n e s f r equen t ly r e q u i r e such a system. The weight increment incurred due t o such a f e a t u r e is included i n t h e weight es t imat ion of t h e f l i g h t con t ro l system, Sect ion 7.1.

6.4.1.1 C w method

For a i r p l a n e s with i n t e r n a l f u e l systems (no t i p t a n k s ) :

wfs 0. 4owF/Kfsp

P a r t V Chapter 6

(6 .15 )

Page 9 0

For a i r p l a n e s wi th external f u e l systems ( w i t h t i p t a n k s ) :

Wfs 0. 70WpIKfsp (6.16)

The cons tant K f s p

takes on t h e fol lowing va lues :

K f s p

= 5.87 l b s l g a l f o r av i a t i on g a s o l i n e

= 6.55 f o r l b s l g a l f o r JP-4

WF = mission f u e l weight ( i n c l u d e s reserves) i n l b s

The f a c t o r Kf is de f ined i n 6.4.1.1.

of new terms:

i n t = f r a c t i o n of f u e l t a n k s which are i n t e g r a l

Nt = number of s e p a r a t e f u e l t a n k s

Ne = number of eng ines

6.4.1.3 T o r w e e k Me-

For t u r b i n e engines , see Sub- section 6.4.2.

For s i n g l e p i s t o n engine i n s t a l l a t i o n s :

Wfs 2(WF/5. 87) 0.667 (6.18)

For m u l t i p i s t o n engine i n s t a l l a t i o n s :

Wfs = 4.5(WF/5. 87) 0.60 (6.19)

6.4.2 C o e T r w o r t Ai-

6.4.2.1 m Methpdl *

For a f u e l system wi th i n t e g r a l tanks see 6.4.2.2.

For a f u e l system wi th s e l f - s e a l i n g b ladder cells:


0 .818 + Wfs = 41.6 ( ( W F / K f s p ) / 1 0 0 ) S UPP (6 .20)

Eor a f u e l s y s t e m w i t h n o n- s e l f - s e a l i n g b l a d d e r cells:

0 .758 + Wfs = 23.1((WF/Kfsp)/100) S UPP (6 .21 )

T h e f a c t o r K f SP

is d e f i n e d i n 6.4.1.1.

W s u ~ ~ is t h e w e i g h t of t h e b l a d d e r support struc-

t u r e and is g i v e n by:

6.4.2.2 To renbeek Method

F o r airplanes e q u i p p e d w i t h n o n- s e l f - s e a l i n g b l a d d e r t a n k s :

F o r a i r p l a n e s e q u i p p e d w i t h i n t e g r a l f u e l t a n k s ( w e t w i n g ) :

6.4.3 M i l - P a w : Bomb and T r a n S p o r t ~imlliuu3S

F o r b a s i c f u e l system w e i g h t s , see S u b- s e c t i o n 6.4.2.

Many mi l i t a ry airplanes carry i n f l i g h t r e f u e l l i n g systems. I n a d d i t i o n , many are e q u i p p e d w i t h f u e l dumping systems. T h e w e i g h t s of t h e s e systems may b e e s t i m a t e d from:

F o r i n - f l i g h t r e f u e l l i n g :

= 1 3 . 6 4 ~ ( W F / K f s p ) / 1 0 0 ~ 0.392 (6 .25) Win f l r e f

F o r f u e l dumping:

Wid. - 7.3 8 ( (WF/Kfsp) 1100) 0 .458 ( 6 . 2 6 )

6.4.4 F i q b t e r & A t t a c k A i m

S e e S u b- s e c t i o n s 6.4.2 and 6.4.3.


6 . 5 PROPULSION SYSTEM YEIGHT ES-

Depending on a i r p l a n e type , t h e p ropu l s ion system weight, W is e i t h e r g iven as a f u n c t i o n of t o t a l eng ine

P weight and lo r miss ion f u e l o r by:

= 'ec + 'ess + Wpc +Wosc, where: ( 6 . 2 2 ) P

Wec = weight of eng ine c o n t r o l s i n l b s

'ess = weight of eng ine s t a r t i n g system i n l b s

W = weight of p r o p e l l e r c o n t r o l s i n l b s PC

Wosc = weight of o i l system and o i l c o o l e r i n l b s

6 . 5 . 1 G- Avia t ion . .

6 . 5 . 1 . 1 C e s n a method

Use a c t u a l da t a .

6 . 5 . 1 . 2 USAF M e w

W is included i n Eqn. ( 6 . 3 1. P

6 .5 .1 .3 Torenbeek Method

W is included i n Eqn. ( 6 . 3 ) . P

6 .5 .2 -1 T r u ~ o r t _ A i w

6 . 5 . 2 . 1 GD M e w

e c o n t r o l s :

For fuselagelwing- root mounted jet engines :

The f a c t o r Ke, takes on t h e fol lowing va lues :

Kec = 0 . 6 8 6 f o r non- afterburning engines = 1 . 0 8 0 f o r a f t e r b u r n i n g engines

For wing mounted jet engines:


For wing mounted turboprops:

-

par wing mounted p i s t o n engines:

Ne = number of eng ines

lf = f u s e l a g e l e n g t h i n f t

b = wing span i n f t

For a i r p l a n e s wi th one o r two jet engines us ing c a r t r i d g e o r pneumatic s t a r t i n g systems:

For a i r p l a n e s wi th fou r o r more jet engines us ing pneumatic s t a r t i n g systems:

For a i r p l a n e s wi th jet engines us ing electric s t a r t i n g systems:

For a i r p l a n e s wi th tu rboprop engines us ing pneumatic s t a r t i n g systems:

For a i r p l a n e s wi th p i s t o n eng ines us ing e l e c t r i c s t a r t i n g systems:

0 .459 'ess = S O . 3 8 ( w e / l 000 )

r c o n t r o l s :

For tu rboprop engines:


For p i s t o n engines:

We = t o t a l weight of a l l eng ines i n l b s

For a i r p l a n e s wi th t u r b o j e t o r t u rbo fan engines us ing c a r t r i d g e o r pneumatic s t a r t i n g systems, t h e weight f o r accessory d r i v e s , powerplant c o n t r o l s , s t a r t i n g and i g n i t i o n ' systems is :

The take- off f u e l f low r a t e , (dWFldt)TO has t h e dimension of l b s l s e c .

For a i r p l a n e s wi th tu rboprop engines t h i s weight is:

The f a c t o r Kb t a k e s on t h e fol lowing va lues :

Kb = 1.0 without b e t a c o n t r o l s = 1.3 with b e t a c o n t r o l s

I t is u s u a l l y accep tab le t o assume t h a t :

- 'api - Wp - 'osc

of new ternrs:

(dWpldt)TO = f u e l f low a t take- off i n l b s l s e c

PTO = requi red take- off power i n hp

u s t r e v e m f o r i e t - m e s :

The weight of t h r u s t r e v e r s e r s was a l r e a d y inc luded i n t h e engine weight e s t i m a t e of Eq. ( 6 . 6 ) . To o b t a i n a b e t t e r estimate of t h e c o g . e f f e c t due t o t h r u s t r e v e r s e r s a s e p a r a t e weight e s t i m a t e is needed:


Water i n j e c t i o n s y s t e m are used t o increase -

take-off performance of a l l types of engines. The i n s t a l l a t i o n of such a system is opt ional .

Wwi = 8. 5 86Wwt ,/ 8.3 5 (6.37)

W w t r = weight of water c a r r i e d i n l b s

Wosc = K~scWe (6.38)

The f a c t o r Kosc takes on t h e following values:

= 0.00 f o r jet engines (weight i n c l . i n We) KOsc = 0.07 f o r turboprop engines

= 0.08 f o r r a d i a l p i s t o n engines = 0.03 f o r h o r i z o n t a l l y opposed

p i s t o n engines

6.5.3 Patrol , Bomb and--

See Sect ion 6.5.2.

6.5.4 Fi-er and A t t a c k A i r g l m s i

See Sect ion 6.5.2.


7 . CLASS I1 METHOD FOR ESTIMATING FIXED EQUIPMENT WEIGHT ---- -,,,p=I=====PS==IPIPS==PPPI=======~=======X===P=====~=~=========

The l is t of f i x e d equipment c a r r i e d on board a i r p l a n e s v a r i e s s i g n i f i c a n t l y wi th a i r p l a n e t y p e and a i r p l a n e mission. I n t h i s chap te r it w i l l be assumed t h a t t h e fo l lowing i tems a r e t o be inc luded i n t h e f i x e d equipment ca tegory :

7 . 1 . F l i g h t c o n t r o l system, W f c

7 .2 . Hydraulic and pneumatic System, W bps

7 .3 . E l e c t r i c a l system, Wels

7 .4 . Ins t rumenta t ion , a v i o n i c s and e l e c t r o n i c s , Wiae

7 .5 . Air- condi t ioning, p r e s s u r i z a t i o n , a n t i - and de- icing system, Wapi

7 .6 . Oxygen system, Wox

7 . 7 . Auxi l i a ry power u n i t (APU), W apu 7.8 . Furn ish ings , W f u r

7 .9 . Baggage and cargo handl ing equipment. Wbc

7 . 1 0 Opera t iona l items, W OPS

7 . 1 1 . Armament, Warm

7 .12 . Guns, l aunchers and weapons p r o v i s i o n s , W glw 7.13 . F l i g h t t es t in s t rumen ta t ion , W f t i

7.14 . Auxi l i a ry g e a r , Waux

7.15 . B a l l a s t , Wbal

7.16 . P a i n t , W p t

7.17 . Wet,

Therefore:


The exact d e f i n i t i o n of which item belongs i n a par- t i c u l a r f ixed equipment category is hard t o f ind . The c a t e g ~ r y Wetc was added t o cover any i tems no t speci-

=#

f i c a l l y l i s t e d .

Methods f o r p red ic t ing weights of t y p i c a l f ixed equipment items a r e presented f o r t h e following types of a i r p l a n e s :

1. General Aviation Airplanes 2. Commercial Transport Airplanes 3. Mi l i t a ry P a t r o l , Bomb and Transport Airplanes 4. F igh te r s and Attack Airplanes

The reader should always consu l t a c t u a l f ixed equipment weight d a t a f o r s i m i l a r a i rp lanes . Appendix A p r e s e n t s t h i s information f o r a l a r g e number of a i rp lanes .

7 . 1 . 1 G e w a l Aviation Ai- . .

where: WTO = take-off weight i n l b s

T h i s equat ion a p p l i e s only t o a i r p l a n e s under 8,000 l b s take-off weight with mechanical f l i g h t con t ro l s . The equat ion inc ludes a l l f l i g h t con t ro l system hardware: cab les , pul leys , pushrods, cockpi t c o n t r o l s p l u s any required back-up s t r u c t u r e .

Airplanes i n t h i s category a l l tend t o have two sets of f l i g h t c o n t r o l s i n t h e cockpit .

For a i r p l a n e s w i t h un-powered f l i g h t con t ro l s :

For a i r p l a n e s with powered f l i g h t con t ro l s :


7.1 .1 .3 Tor-

For a i r p l a n e s with un-powered, unduplicated f l i g h t con t ro l s :

The following equat ion a p p l i e s t o business jets a s w e l l a s t o commercial t r a n s p o r t a i rp lanes :

- where: qD is t h e design d ive dynamic p ressu re i n psf

The cons tant K f c t a k e s on t h e following values:

= 0 .44 f o r a i r p l a n e s with un-powered f l i g h t 'fc c o n t r o l s

= 0.64 f o r a i r p l a n e s with powered f l i g h t c o n t r o l s

I f leading edge devices a r e employed, t h e s e es t imates should be mul t ip l ied by a f a c t o r 1.2. I f l i f t dumpers a r e employed, a f a c t o r 1.15 should be used.

where: < is t h e design d ive dynamic p ressu re i n psf L

where: Ssc is t h e t o t a l c o n t r o l s u r f a c e a r e a i n f t 2

mte: t h e s e e s t ima tes include t h e weight of a l l assoc ia ted hydraul ic andlor pneumatic systems1


For USAF f i g h t e r s :

The cons tant K f c f t a k e s on t h e following values:

= 1 0 6 f o r a i r p l a n e s with elevon c o n t r o l Kfcf and no hor izon ta l t a i l

= 138 f o r a i r p l a n e s with a hor i zon ta l t a i l

= 1 6 8 f o r a i r p l a n e s with a v a r i a b l e sweep wing

For USN f i g h t e r s and a t t a c k a i r p l a n e s :

Note: t h e s e es t imates include t h e weight of a l l a s soc ia ted hydraul ic andlor pneumatic systems.

Cer ta in a i r p l a n e s r e q u i r e a c e n t e r of g r a v i t y c o n t r o l system. T h i s is normally implemented using a f u e l t r a n s f e r system. The e x t r a weight due t o a c.g. con t ro l system may be est imated from:

4

where: WF is t h e mission f u e l weight i n l b s

K f SP

= 6 . 5 5 l b s / g a l f o r JP- 4


7.2 HYDRAULIC W I O R P W - E M -HI' ES-

A s seen i n Sect ion 7 .1 t h e weight of t h e hydraul ic and/or pneumatic system needed f o r powered f l i g h t c o n t r o l s is usual ly included i n t h e f l i g h t c o n t r o l system weight p red ic t ion .

The following weight r a t i o s may be used t o determine t h e W i c svst- we- s epa ra te ly :

For business j e t s : 0.0070 - 0.0150 of WTO

For regional turboprops: 0.0060 - 0.0120 of WTO

For commercial t r a n s p o r t s : 0.0060 - 0.0120 of WTO

For m i l i t a r y p a t r o l , t r a n s p o r t and bombers:

For f i g h t e r s and a t t a c k a i rp lanes :

The reader should consul t t h e d e t a i l e d weight d a t a i n Appendix A f o r more p r e c i s e information.

J . 3 ELEeTRICAt SYSTEM U G H T ES-

The reader should consul t t h e d e t a i l e d weight d a t a i n Appendix A f o r e l e c t r i c a l system weights of s p e c i f i c a i rp lanes .

Note t h a t t h e e l e c t r i c a l system weight i n t h i s case is given a s a funct ion of t h e weight of t h e f u e l system p l u s t h e weight of instrumentat ion, av ion ics and e l e c t r o n i c s . t

P a r t V Chapter 7 Page 1 0 1

'where: WE is t h e empty weight i n l b s

where: V i s t h e passenger cabin volume i n it 3 Pax

7 . 3 . 3 WV - P a u o l . Bomb T r m ~ o r t Airr>lanes

7 . 3 . 3 . 1 GD M e u

F o r t r a n s ~ o r t airDlanes:

Use Eqn. ( 7 . 1 5 )

7 . 3 . 4 . 1 W M e w

For USAF f i g h t e r s :

For USN f i g h t e r s and at tack a i rp lanes :

part v Chapter 7 Page 102

The reader should c o n s u l t t h e d e t a i l e d weight d a t a i n Appendix A f o r weights of i n s t rumen ta t ion , a v i o n i c s and e l e c t r o n i c s f o r s p e c i f i c a i r p l a n e s . Another important source of weight d a t a on a c t u a l a v i o n i c s and e l e c t r o n i c s systems f o r c i v i l a i r p l a n e s is Reference 18. For d a t a on m i l i t a r y a v i o n i c s systems t h e reader should c o n s u l t Reference 1 3 , Tables 8-1 and 8-2.

o r t a n t c o m The weight e q u a t i o n s g iven i n t h i s s e c t i o n a r e o b s o l e t e f o r modern EFIS t y p e c o c k p i t i n s t a l l a t i o n s and f o r modern computer based f l i g h t management and nav iga t ion systems. The equa t ions provided a r e probably conserva t ive .

7.4.1 G-al Avia t ion AirDlanes . .

where: N Pax

is t h e number of passengers , i nc lu- ding t h e crew

For t h e weight of ins t ruments :

'i P

Npil 1 + 0.032(WT0/1,000)I + N e f S + 0*006(WT0/1D000)I +

f l i g h t ins t ruments eng ine in s t rumen t s

+ 0.15(wTo11,000) + 0.012WT0

o t h e r i n s t rumen t s

where: N p i 1

is t h e number of p i l o t s

We is t h e number of eng ines ,


7.4 .2 .2 T o r e n b e e k M e u

'iae = 1 2 0 + 20Ne + 0.006WT0

For i e t tr-

w h e r e : WE is t h e empty w e i g h t i n l b s

R is t h e maximum range i n naut ical miles

7 .4 .3 W v P w , Bomb a n d Tr-t A i n > l a n e s

U s e S u b- s e c t i o n 7.4.2.

7.4.4 F i a h t e r and A t t a c k

Use S u b- s e c t i o n 7.4.2.

7 .5 WEIGHT ESTIMATION FOR A I R I T I O N I N G , - G SYSTEMS

where: N Pax is t h e number of passengers, i n c l u d i n g

t h e crew MD is t h e d e s i g n d i v e Mach number

P a r t V C h a p t e r 7 P a g e 1 0 4

7 . 5 . 2 C o m a 1 Tra-ort

7 . 5 . 2 . 1 GD Me-

For - a i rp lanes :

7 . 5 . 2 . 2 Tor-

For - a i rp lanes :

where 1 Pax

= l eng th of t h e passenger cabin i n f t

7 . 5 . 3 W v - Patyol , Bomb a Trans-

The constant Kapi t a k e s on t h e following values:

= 887 f o r subsonic a i r p l a n e s with wing and t a i l ',pi an t i- ic ing

= 610 f o r subsonic a i r p l a n e s without an t i- ic ing

= 7 4 8 f o r supersonic a i r p l a n e s without an t i- ic ing

7 . 5 . 4 . 1 GD Method

For low subsonic W e s r

The cons tant K a p i

t a k e s on t h e following values:

= 2 1 2 f o r a i r p l a n e s with wing and t a i l p i an t i- ic ing

= 109 f o r a i r p l a n e s without an t i- ic ing *


7 . 6 . 1 G w Avlat lon . . 4

Use Sub- section 7 . 6 . 2 .

m t e : Equation number ( 7 . 3 4 ) has been i n t e n t i o n a l l y de l e t ed .

For commercial transport a i rp lanes and f o r bus ines s type a i rp lanes :

For f-ts below 25.000 fU

For s h o r t f w above 25.000 f t :

For cued overwater fliahts:

7 . 6 . 3 W v Patro l . Romb rind Tr-t

U s e Sub- section 7 . 6 . 2 .

Part V Chapter 7 Page 1 0 6

Auxi l i a ry power u n i t s a r e o f t e n used i n t r a n s p o r t o r p a t r o l t y p e a i r p l a n e s , commercial as we l l a s m i l i t a r y .

Actual APU manufacturer d a t a should be used, where poss ib l e . Reference 8 c o n t a i n s d a t a on APU systems, under 'Engines ' .

From t h e d e t a i l e d weight s t a t e m e n t s i n Appendix A it is p o s s i b l e t o d e r i v e weight f r a c t i o n s f o r t h e s e systems as a f u n c t i o n of t h e take- off weight, WTO. The fol lowing

ranges a r e t y p i c a l of t h e s e weight f r a c t i o n s :

The f u r n i s h i n g s ca t ego ry normal ly i n c l u d e s t h e fo l lowing items:

1. s e a t s , i n s u l a t i o n , t r i m pane l s , sound proof ing , ins t rument pane l s , c o n t r o l s t a n d s , l i g h t i n g and wi r ing

2. Ga l ley ( p a n t r y ) s t r u c t u r e and p r o v i s i o n s

3. Lavatory ( t o i l e t ) and a s s o c i a t e d systems

4. Overhead luggage c o n t a i n e r s , h a t r a c k s , wardrobes

5. Escape p rov i s ions , f i r e f i g h t i n g equipment

Note: t h e a s s o c i a t e d consumable i tems such as po- t a b l e water, food, beverages and t o i l e t chemicals and pa- p e r s are normally inc luded i n a weight ca tegory r e f e r r e d t o as: Opera t iona l I tems: Wops, see S e c t i o n 7.10.

The r eade r is r e f e r r e d t o t h e d e t a i l weight s t a t emen t s i n Appendix A f o r a c t u a l f u r n i s h i n g s weight d a t a on s p e c i f i c a i r p l a n e s .

where: N Pax

is t h e number of passengers i nc lud ing t h e crew


7.8.1.2 Tor- Method:

For s i n g l e engine a i rp lanes : -3

'fur = 5 + 1 3 N p a x + 25NrOw,

where: Nrow is t h e number of s e a t rows

For mult i engine a i r p l a n e s :

W i u r a lSNpax + l*ovpax+cargop ( 7 . 4 3 )

Where: Vpax+cargo is t h e volume of t h e passenger

cabin p l u s t h e cargo volume i n f t 3

The weight of fu rn i sh ings v a r i e s considerably with a i r p l a n e type and with a i r p l a n e mission. This weight i t e m is a considerable f r a c t i o n of t h e take-off weight of most a i r p l a n e s , a s t h e d a t a i n Appendix A i l l u s t r a t e .

Reference 1 4 con ta ins a very d e t a i l e d method f o r est imating t h e fu rn i sh ings weight f o r commercial t r a n s p o r t a i rp lanes .

7. 8.2.1 GD Method

f d c sts pax sts c c sts l a v s + water food prov.

+ 1 0 9 ( ( N (1 + PC)/lOO) O o 5 0 5 + 0. 771(WTo/1.000) Pax

cabin windows miscellaneous

The f a c t o r Klav t akes on t h e following values:

= 3 . 9 0 f o r bus iness a i r p l a n e s K1av = 0.31 f o r s h o r t range a i r p l a n e s

= 1.11 f o r long range a i r p l a n e s

The f a c t o r Kbuf t a k e s on t h e following values:

= 1 .02 f o r s h o r t ranges Kbuf = 5.68 f o r very long ranges


T h e t e r m PC is t h e des ign u l t i m a t e cab in p r e s s u r e

i n p s i . The va lue of PC depends on t h e des ign a l t i t u d e

f o r t h e p r e s s u r e cabin.

I n commercial t r a n s p o r t s it is u s u a l l y d e s i r a b l e t o make more d e t a i l e d estimates t h a n p o s s i b l e wi th Eqn. ( 7 . 4 5 ) . P a r t i c u l a r l y i f a more a c c u r a t e l o c a t i o n of t h e c o g . of items which c o n t r i b u t e t o t h e f u r n i s h i n g s weight is needed, a more d e t a i l e d method may be needed. Reference 1 4 c o n t a i n s t h e necessary d e t a i l e d informat ion.

7.8.3 m v P a t r o l . B o m b a n d o r t

'fur = Sum i i n t h e t a b u l a t i o n below. ( 7 . 4 6 )

TYPe P a t r o l Bomb Transpor t

1 .2 1 . 2 Crew E j . K s t ( N c r ) S e a t s K s t ( N c r )

= 1 4 9 w i th s u r v i v a l k i t Kst = 100 wi thout s u r v i v a l k i t

Crew S e a t s 8 3 ( ~ ~ ~ ) ~ * ~ ~ ~ same same

Passenger S e a t s

Troop S e a t s

Lav. and Water

0. 0019 twTo) 0. 839 Misc. 0. 771(WT,/l,000)

e j e c t i o n seats Misc. and emergency eqpmt


%he GD method g i v e s f o r m i l i t a r y passenger t r a n s p o r t s :

The c o n s t a n t Kbc t a k e s on t h e fo l lowing va lues :

Kbc = 0.0646 wi thout p re load p r o v i s i o n s = 0.316 wi th pre load p r o v i s i o n s

The Torenbeek method g i v e s f o r commercial ca rgo a i r p l a n e s :

'bc = 3 S f f n ( 7 . 4 9 )

where: 2 Sff is t h e f r e i g h t f l o o r a r e a i n f t . For baggage and f o r ca rgo c o n t a i n e r s , t h e fol lowing

weight e s t i m a t e s may be used:

f r e i g h t p a l l e t s : 88x108 i n ( i nc lud ing n e t s ) 88x125 i n

96x125 i n

2 2 5 l b s 2 6 2 l b s 2 8 5 l b s

c o n t a i n e r s : 1 . 6 l b s / f t 3 (For c o n t a i n e r dimensions, s e e P a r t 111 .1

Typica l weights counted i n o p e r a t i o n a l i tems a r e :

*Food *Potab le water *Drinks

*China *Lavatory s u p p l i e s

Observe t h a t Eqn. ( 7 . 4 4 ) i n c l u d e s t h e s e o p e r a t i o n a l i tems. For more d e t a i l e d in format ion on o p e r a t i o n a l i tems t h e reader should c o n s u l t Reference 1 4 , p.292.

7 . 1 1 ARMAMENT WEIGHT

The ca t ego ry armament can c o n t a i n a wide v a r i e t y of weapons r e l a t e d i tems a s w e l l a s p r o t e c t i v e s h i e l d i n g f o r t h e crew. Typica l armament items a r e :


*F i r ing systems * F i r e c o n t r o l systems

*Bomb bay o r missile doors *Armr p l a t i n g

*Weapons e j e c t i o n systems

Note t h a t t h e weapons themselves as w e l l as any ammunition a r e n o t normally inc luded i n t h i s item.

Appendix A c o n t a i n s d a t a on 'armament' weight f o r s e v e r a l t y p e s of m i l i t a r y a i r p l a n e s .

7.12 W U H T -N FOR GUNS. - For d e t a i l e d d a t a on guns, l a u c h e r s and o t h e r

m i l i t a r y weapons p r o v i s i o n s t h e r eade r is r e f e r r e d t o P a r t 111, Chapter 7.

Note: Ammunition, bombs, missiles, and most t y p e s of e x t e r n a l s t o r e s a r e normally counted as p a r t of t h e payload weight, Wpl i n m i l i t a r y a i r p l a n e s .

During t h e c e r t i f i c a t i o n phase of most a i r p l a n e s a s i g n i f i c a n t amount of f l i g h t t es t in s t rumen ta t ion and a s s o c i a t e d hardware is c a r r i e d on board. The magnitude of W f t i depends on t h e t ype of a i r p l a n e and t h e t y p e s of

f l i g h t tes ts t o be performed. Appendix A c o n t a i n s weight d a t a f o r f l i g h t test in s t rumen ta t ion c a r r i e d on a number of NASA exper imental a i r p l a n e s (Tables A13.1-A13.4).

7.14 -N FOR

Th i s item encompasses such equipment as:

* f i r e axes * s e x t a n t s *unaccounted i tems

An i t e m r e f e r r e d t o a s 'manufacturers v a r i a t i o n ' is sometimes inc luded i n t h i s ca t ego ry as w e l l . A safe assumption is t o set:

When looking over t h e weight s t a t e m e n t s f o r v a r i o u s a i r p l a n e s i n Appendix A, t h e r eade r w i l l make t h e s t a r t l i n g d i scovery t h a t some a i r p l a n e s c a r r y a


s i g n i f i c a n t amount of b a l l a s t . This can have d e t r i m e n t a l e f f e c t s on speed, payload and range performance.

The fol lowing reasons can be g iven f o r t h e need t o i nc lude b a l l a s t i n an a i r p l a n e :

1. The des igne r 'goofed' i n t h e weight and ba lance c a l c u l a t i o n s

2. To ach ieve c e r t a i n aerodynamic advantages it was judged necessary t o l o c a t e t h e wing o r t o s i z e t h e empennage s o t h a t t h e s t a t i c margin became i n s u f f i c i e n t . T h i s problem can be so lved wi th b a l l a s t . I n t h i s case, c a r r y i n g b a l l a s t may i n f a c t t u r n o u t t o be advantageous.

3. To ach ieve f l u t t e r s t a b i l i t y w i th in t h e f l i g h t envelope b a l l a s t weights a r e sometimes a t t a c h e d t o t h e wing and/or t o t h e empennage.

Note: ba lance weights a s s o c i a t e d wi th f l i g h t c o n t r o l s u r f a c e s a r e n o t counted a s b a l l a s t weight.

The amount of b a l l a s t weight r equ i r ed is determined wi th t h e h e l p of t h e X-plot. Cons t ruc t ion and u s e of t h e X-plot is d i scussed i n P a r t 11, Chapter 11. The C l a s s I1 weight and ba lance method d i scussed i n Chapter 9 of t h i s p a r t may a l s o be h e l p f u l i n determining t h e amount of b a l l a s t weight r equ i r ed t o ach ieve a c e r t a i n amount of s t a t i c margin.

T ranspor t jets and camouflaged m i l i t a r y a i r p l a n e s c a r r y a cons ide rab le amount of p a i n t . The amount of p a i n t weight is obvious ly a f u n c t i o n of t h e e x t e n t of s u r f a c e coverage. For a w e l l p a i n t e d a i r p l a n e a reasonable estimate f o r t h e weight of p a i n t is:

This weight i t em has been inc luded t o cover any items which do n o t normally f i t i n any of t h e prev ious weight c a t e g o r i e s .


8. LOCATING COMPONENT CENTERS OF GRAVITY = = S S P I 3 1 P I I P I P = P e l P = f = e ; = = = I t = s = = 6 5 s P P P 6 L :

The purpose of t h i s chapter is t o provide gu ide l ines f o r t h e determination of t h e l o c a t i o n of cen te r s of g r a v i t y f o r ind iv idua l a i r p l a n e components. Knowledge of component cog . l o c a t i o n s is e s s e n t i a l i n both Class I and Class I1 weight and balance analyses a s discussed i n Chapter 1 0 of P a r t I1 and Chapter 4 of t h i s book.

In P a r t 11, Chapter l o , Table 10.2 provides a summary of c o g . l o c a t i o n s f o r t h e major s t r u c t u r a l components of t h e a i r p l a n e only. In t h i s chapter a s l i g h t l y more extens ive d a t a base is provided. The p resen ta t ion of component c o g . l o c a t i o n s follows t h e weight breakdowns of Chapters 5-7:

8.1 C.G. Locations of S t r u c t u r a l Components 8.2 C.G. Locations of Powerplant Components 8.3 C.G. Locations f o r Fixed Equipment

Table 8.1 l ists t h e most l i k e l y c.g. l o c a t i o n s f o r major s t r u c t u r a l components. There i s no s u b s t i t u t e f o r common sense: i f t h e prel iminary s t r u c t u r a l arrangement of P a r t 111 (Step 1 9 of p.d. sequence 2 , P a r t 11) sugges ts t h a t a given s t r u c t u r a l component has a d i f f e r e n t mass d i s t r i b u t i o n than is commonly t h e case , an 'educated guess ' must be made a s t o t h e e f f e c t on t h e c.g. of t h a t component.

ExamDle: Looking a t t h e threeview of t h e GP-180 of Figure 3 .47 , p.86, P a r t I1 it is obvious t h a t t h e r e is a concent ra t ion of primary s t r u c t u r e a t t h e a f t end of t h e fuselage. The fuse lage c o g . should t h e r e f o r e no t be placed a t 38-40 percent of t h e fuse lage length , but probably a t 55 t o 60 percent .

8.2 C.G. L- OF PO- C-

Table 8.2 l ists t h e most l i k e l y cog . l o c a t i o n s f o r powerplant components. Note t h a t f o r engine c.g. l o c a t i o n s manufacturers d a t a should be used. 'Guessing' a t engine cog . l o c a t i o n s is not recommended!

8.3 C.G. L-NS OF F U P D R m

Table 8.3 lists gu ide l ines f o r loca t ing cen te r s of g r a v i t y of f ixed equipment components.

Part V Chapter 8 Page 113

Table 8 .1 Center of Grav i ty Locat ion of S t r u c t u r a l Components ~ P P l P = = ~ P P P P P P P I P ~ P P L i P P P I ~ P P P m m ~ E n ~ ~ P P P P ~ ~ ~ n ~ ~ P ~ ~ a l ~ n n u n ~ a ~ a ~

Component :

Wing ( h a l f ) :

Horizontal T a i l : ( h a l f V e r t i c a l T a i l : (low t a i l ) V e r t i c a l T a i l : ( T - t a i l ) r oo t V e r t i c a l T a i l : ( c ruc i form)

Fuselage :

Caution: Do no t count t h e p r o p e l l e r sp inne r i n f u s e l a g e o r n a c e l l e l eng th!

T a i l booms:

Nacelles :

Landing gea r :

Center of g r a v i t y l o c a t i o n : 4

w i a 38-42 p e r c e n t chord from t h e L.E. a t 40 pe rcen t of t h e semi-span.

wina: 70 p e r c e n t of t h e d i s t a n c e between t h e f r o n t and r e a r s p a r behind t h e f r o n t s p a r a t 35 pe rcen t of t h e semi-span

Regardless of sweep angle : 4 2 p e r c e n t chord from t h e L. E. a t 38 pe rcen t of t h e semi-span. Regardless of sweep angle : 4 2 p e r c e n t chord from t h e L.E. a t 38 pe rcen t v e r t i c a l t a i l span from t h e r o o t chord. Regardless of sweep angle: 4 2 p e r c e n t chord from t h e L.E. a t 55 pe rcen t v e r t i c a l t a i l span from t h e r o o t chord. Regardless of sweep angle : 4 2 p e r c e n t chord from t h e L.E. a t between 38 and 5 5 pe rcen t v e r t i c a l t a i l span from t h e r o o t chord. I n t e r p o l a t e according t o zh/bv.

S i n g l e engine t r a c t o r s : 0.32-0.35 S i n g l e engine pusher: 0.45- 0.4 8 P r o p e l l e r d r i v e n twins : 0.38-0.40 ( t r a c t o r s on wing) P r o p e l l e r d r i v e n twins: 0.50-0.53 (pushers on wing J e t t r a n s p o r t s : 0.42-0.45 (wing mounted engines) Jet t r a n s p o r t s : 0.47-0.50 (rear f u s e l a g e mounted engines) F i g h t e r s : 0 .45 (engines bu r i ed i n t h e f u s e l a g e )

0.40-0.45 of boom l e n g t h s t a r t i n g from most forward s t r u c- t u r a l a t tachment of t h e boom-

0 .40 of n a c e l l e l e n g t h from n a c e l l e nose

a t 0.50 of s t r u t l e n g t h f o r g e a r s wi th mostly v e r t i c a l s t r u t s

Table 8.2 Center of Grav i ty Locat ion of Powerplant = = = = P = e P P = ~ a a P I P P P I I P t ~ m P P n a = = ~ P a = ~ ~ a ~ P P 6 E ~ ~ m ~ ~ P ~ ~

Components a = = P D L P P t P

Component: Center of Grav i ty Location:

Engine(s ) Use manufacturers d a t a

Air i nduc t i o n s y s tem Use t h e c.g. of t h e g r o s s s h e l l a r e a of t h e i n l e t s

P r o p e l l e r s

Fuel system

On t h e s p i n a x i s , i n t h e pro- p e l l e r s p i n p l a n e

Refer t o t h e f u e l system l ayou t diagram requ i r ed as p a r t of S t ep 1 7 i n p o d . sequence 11, P a r t 11, p. 18.

F i l l e d f u e l t ank Assuming a pr i smoida l shape (See f i g u r e l e f t ) , t h e c.g. is loca t ed r e l a t i v e t o p l a n e S1 a t :

( 1 1 4 ) IS1 + 3s2 + ~ ( s ~ s ~ ) ~ ~ ~ I / { s ~ + S2 +(S1S2) 1 / 2 ] leg

Trapped f u e l and o i l Trapped f u e l is normally l o- ca t ed a t t h e bottom of f u e l t a n k s and f u e l l i n e s . Trapped o i l is normally l o- c a t e d c l o s e t o t h e engine case .

Propuls ion system

P a r t V

Make a list of which i tems c o n t r i b u t e t o t h e p ropu l s ion system weight and 'gues t imate ' t h e i r c.g. l o c a t i o n by r e f e r r i n g t o t h e powerplant i n s t a l l a t i o n drawing r equ i r ed i n S t e p 5.10, pages 1 3 3 and 1 3 4 i n P a r t 11.

Chapter 8 Page 1 1 5

Table 8.3 Center of Gravi ty Location of Fixed Equipment = = ' = S P S t l a P P = D I D I ~ I P 3 1 I P ~ P P ~ ~ ~ ~ ~ ~ ~ ~ 3 1 I u % P P t ~ = P ~ = 5 a ~ ~ ~ ~ ~ ~ ~

Component:

F l i g h t Control System

Hydraulic and Pneumatic System

E l e c t r i c a l System

Instrumentat ion, Avionics and E lec t ron ics

Center of Gravity Location:

mte: f o r a l l systems, t h e cog . l o c a t i o n can be most c l o s e l y *guest imatedt by r e f e r r i n g t o t h e system lay-

b ou t diagrams descr ibed i n P a r t I V of t h i s t e x t , These system lay- outs were required a s p a r t of S tep 1 7 i n p.d. sequence 2 , P a r t 11, p.18.

Air- conditioning, Pressu- r i z a t i o n , Anti- icing and de- icing System

Oxygen System J Auxil iary Power Unit See engine manufacturer da ta ,

Furnishings Refer t o t h e fuse lage i n t e r - n a l arrangement drawing re-

Baggage and Cargo qui red by Steps 4 . 1 and 4.2 Equipment i n P a r t 11, pp 1 0 7 and 108.

Operational i tems See fu rn i sh ings

Armament This i t e m is normally c l o s e t o t h e cockpi t

Guns, launchers and wea- From manufacturer data . pons provis ions

F l i g h t t e s t i n s t r u - A ske tch depic t ing t h e loca- mentat ion t i o n s of sensors , recorders

opera t ing systems should he lp i n loca t ing t h e o v e r a l l c o g . of t h i s item.

Auxil iary gear Make a l i s t of items i n t h i s category and 'guestimate ' t h e i r cog . loca t ions .

B a l l a s t B a l l a s t weights a r e normally made from lead. B a l l a s t c.g. is thus e a s i l y located.

Pa in t

P a r t V

Centroid of painted areas .

Chapter 8 Page 1 1 6

9. CLASS I1 WEIGHT AND BALANCE ANALYSIS PP==BP===IPP=PPIDPP=5I5==r: I=trSPPii:PI====

The b a s i c method used i n performing a Class I1 weight and balance a n a l y s i s is i d e n t i c a l t o t h a t used f o r t h e C l a s s I weight and balance a n a l y s i s . The l a t t e r was d i scussed i n d e t a i l i n P a r t 11, Chapter 10. The on ly d i f f e r e n c e is, t h a t a more d e t a i l e d weight s t a t emen t is used: t h e Class I1 weight p r e d i c t i o n method of Chapters 4-7 is used.

During t h i s s t a g e of t h e p re l imina ry des ign f r e q u e n t q u e s t i o n s which a r e r a i s e d , a r e :

1. How much does t h e o v e r a l l a i r p l a n e c.g. move as a r e s u l t of moving some component?

2. How much does t h e a i r p l a n e s t a t i c margin change as a r e s u l t of moving t h e wing?

These q u e s t i o n s a r e answered i n S e c t i o n s 9 . 1 and 9.2 r e s p e c t i v e l y .

F igu re 9 . 1 i l l u s t r a t e s an a i r p l a n e , i t s c.g. l o c a t i o n and t h e c.g. l o c a t i o n of component i. The o v e r a l l c e n t e r of g r a v i t y of t h e a i r p l a n e is found from:

x - (Sum Wixi / (Sum Wi) c g is1

Evident ly:

i = n Sum Wi = W i-1

The r a t e a t which o v e r a l l a i r p l a n e c.g. moves, when a component i is moved, can be found by d i f f e r e n t i a t i o n of Eqn. (9.1):

i = n ax laxi - (wi) 1 (Sum Wi)

Cg i=l

If component i is moved over a d i s t a n c e Axi, t h e


i

t

- X - - m e 9 . 1 Deflnlt lon of Over& C . G . LoQ3,iion U l d of . . .

G. Location

NOTE: fAnT XI Y CONTAINS A METHOD

t FDA T CONSTRUCTION

Z

Part V Chapter 9 Page 1 1 8

o v e r a l l a i r p l a n e c.g. moves over a d i s t a n c e g iven by:

i = n Ax = (Axi 1 (Wi 1 (Sum Wi

cg ill

Equation ( 9 . 4 ) sugges t s t h a t t o move t h e o v e r a l l c.g. of t h e a i r p l a n e s i g n i f i c a n t l y , e i t h e r a heavy weight component can be moved a small d i s t a n c e o r a l i g h t weight component can be moved a l a r g e d i s t a n c e .

Items which a r e f r e q u e n t l y moved about t o ach ieve s a t i s f a c t o r y weight and ba lance r e s u l t s a r e : b a t t e r i e s , a i r- c o n d i t i o n e r u n i t s * c e r t a i n 'b lack ' boxes and sometimes j u s t p l a i n b a l l a s t . The r eade r w i l l n o t e from t h e d e t a i l e d weight s t a t emen t s i n Appendix A t h a t s e v e r a l a i r p l a n e s c a r r y a r e l a t i v e l y l a r g e amount of b a l l a s t .

9.2 EFFECT OF M O V I N G ~ L ~ ITY BbID ON O V V C GENU3.B

Figure 9.2 i l l u s t r a t e s an a i r p l a n e , i ts mean geometr ic chord l o c a t i o n and its o v e r a l l c.g. l o c a t i o n .

I f t h e lead ing edge of t h e mgc of t h e wing is a t fu- s e l a g e s t a t i o n (PSI XLE, t h e a i r p l a n e c.g. i n t e r m of

t h e wing mgc can be w r i t t e n as: - x - (xcg - xLE) 1; ( 9 . 5 ) cg When a component i is moved over a d i s t a n c e Axi,

t h e o v e r a l l a i r p l a n e c.g. moves r e l a t i v e t o t h e wing mgc as :

- - i-n Ax - (Axi)(Wi)/ctSumWi)

Cg ill

For a convent iona l , t a i l - a f t a i r p l a n e , i ts aerodynamic c e n t e r l o c a t i o n can be w r i t t e n as:

where: C1 = x + Ax acw aCwb

C2 n (CL ICL 1 (1 - d s l d a ) (Sh/S) (9 .9 )

=h a wb P a r t V Chapter 9 Page 1 1 9

A d e t a i l e d d e r i v a t i o n of Eqn. ( 9 . 7 ) may be found i n Reference 1 9 , p.133.

P a r t VI c o n t a i n s methods f o r computing t h e l i f t c u r v e s l o p e s and aerodynamic c e n t e r s which appear i n C1 and i n C2*

t h e wing-body aerodynamic c e n t e r s h i f t , -

Ax i n Eqn. (9 . 8 ) is always a nega t ive number: it "wb

s h i f t s t h e a.c. forward!

I f t h e wing is moved a f t over a d i s t a n c e Axw, t h e o v e r a l l a i r p l a n e c.g. is:

- - - i = n x = x + (AxwWw) 1 sum Wi) ( 9 .10 )

'gnew i-1

The new a.c. l o c a t i o n can now be w r i t t e n a s :

Equat ions ( 9 . 1 0 ) and ( 9 . 1 1 ) can be used t o ' redo ' t h e X-plot of P a r t 11, Chapter 11. Th i s ' redone ' X-plot i n t u r n is used t o :

1. determine how much t h e h o r i z o n t a l t a i l a r e a must be changed a s a r e s u l t of moving t h e wing

o r :

2. determine which o t h e r weight components need t o be moved and by how much, t o main ta in some d e s i r e d l e v e l of s t a b i l i t y ( o r i n s t a b i l i t y a s t h e c a s e may be) .

For a canard a i r p l a n e and /o r f o r a t h r e e s u r f a c e a i r p l a n e similar equa t ions are e a s i l y der ived . The r eade r should c o n s u l t Equat ions (11.1) and ( 1 1 . 2 ) i n P a r t I1 f o r guidance.


10. CLASS I1 METHOD FOR ESTIMATING AIRPLANE INERTIAS 3PP==IPPPPPPPPP6IXP=%=P=P=============PP=z==S====z=z

The purpose of t h i s chap te r is t o prov ide an o u t l i n e f o r a Class I1 method f o r e s t ima t ing moments and produc ts of i n e r t i a . I t w i l l be assumed t h a t t h e C l a s s I1 weight e s t ima t ing method of Chapter 4 has been app l i ed : a r a t h e r d e t a i l e d weight and c.g. breakdown f o r t h e a i r p l a n e is t h e r e f o r e presumed t o be a v a i l a b l e .

The fol lowing equa t ions a r e a s l i g h t mod i f i ca t ion of t h e g e n e r a l i n e r t i a equa t ions 2 . 2 2 a through 2 . 2 2 ~ i n Reference 1 9 .

i =n I,, = Sum mi{ (yi - Ycg l 2 + (z i - z 121

i=1 Cg

i-n = Sum mi((zi - z j 2 + ( X - x ) 2 ~

IYY i-1 cg i cg

i = n 2 I,, - Sum mi((xi - x + (yi -

is1 Cg

i =n = Sum mi(yi -

Ycg ) ( z i - z ) IYZ i-1 cg

F igu re 10.1 d e f i n e s t h e c o o r d i n a t e s used i n t h e s e equa t ions .

The reader should recall t h a t f o r a symmetrical a i r p l a n e t h e i n e r t i a p roduc ts I and I a r e zero.

XY Y Z

Equat ions (10.1) through (10 .6) a r e v a l i d whenever t h e weight breakdown c o n t a i n s a ' s u f f i c i e n t l y g large number of p a r t s s o t h a t t h e i n e r t i a moment and lo r product of each p a r t about its own c.g. l o c a t i o n is n e g l i g i b l e .

Whenever t h e l a t t e r assumption is n o t s a t i s f i e d ,

P a r t V Chapter 1 0 Page 1 2 1

equa t ions (10.1) through (10 .6 ) should be a l l modified as fo l lows :

i = n i-n *

*xx = sum *xxi + Sum mi( (yi - ycg )' + ( z i - z )21 ( 1 0 . 7 ) i-1 i51 cg

The f i r s t t e rm i n Eqn. ( 1 0 . 7 ) r e p r e s e n t s t h e moment ( o r p roduc t ) of i n e r t i a of component i about its own c e n t e r of g r a v i t y .

Moments (and produc ts ) of i n e r t i a of a i r p l a n e components about t h e i r own c e n t e r of g r a v i t y can b e computed i n a r e l a t i v e l y s t r a i g h t f o r w a r d manner by assuming uniform mass d i s t r i b u t i o n s f o r s t r u c t u r a l components and by using t h e 'lumped massp assumption f o r d i s t r i b u t e d systems. An example of t h e l a t t e r would be t h e a i r p l a n e f u e l system. Major f u e l system components such as pumps. b l adde r s and t h e l i k e can be cons idered t o be concent ra ted masses d i s t r i b u t e d around t h e f u e l system c.g.. Equations (10 .1) through (10.2) a r e t hen used t o compute t h e moments of i n e r t i a of t h e f u e l system about i t s own c.g.

K

P a r t V Chapter 1 0 Page 1 2 2

11. REFERENCES PPIP='IPPPIPP=

1. Roskam, J.. Ai rp lane Design: P a r t I, Pre l iminary S i z ing of Airplanes .

2. Roskam, J., Ai rp l ane Design: P a r t 11, Pre l iminary Conf igura t ion Design and I n t e g r a t i o n of t h e Propuls ion System.

3. Roskam, J.. Ai rp lane Design: P a r t 111, Layout Design of Cockpi t , Fuselage, Wing and Empennage: Cutaways and Inboard P r o f i l e s .

4. Roskam, J., Ai rp lane Design: P a r t I V , Layout Design of Landing Gear and Systems.

5. Roskam, J., Ai rp lane Design: P a r t V I , P re l iminary C a l c u l a t i o n of Aerodynamic, Thrus t and Power C h a r a c t e r i s t i c s .

6. Roskam, J.. Ai rp lane Design: P a r t V I I , Determinat ion of S t a b i l i t y , Cont ro l and Performance C h a r a c t e r i s t i c s : FAR and M i l i t a r y Requirements.

7. Roskam, J., Ai rp lane Design: P a r t V I I I , Ai rp lane Cost Es t imat ion and Opt imizat ion: Design, Development Manufacturing and Operating.

Note: These books are a l l publ i shed by: Roskam Avia t ion and Engineering Corpora t ion , Rt4, Box 274, O t t a w a , Kansas, 66067, Tel . 913-2421624.

8. Taylor , J.W.R., J a n e ' s A l l The World A i r c r a f t , Publ ished Annually by: Jane's pub l i sh ing Company, 238 C i t y Road, London EClV 2PU, England. (Issues used: 1945146, 1968184)

9. Chawla, J.P., Empir ical Formulae f o r Rad i i of Gyra- t i o n of A i r c r a f t , SAWE Paper No.78, Hughes A i r c r a f t Company, Culver C i t y , C a l i f o r n i a , 1952.

10. Anon., Empirical Formulae f o r Moments of I n e r t i a of A i r c r a f t , Royal A i r c r a f t Establ ishment , S t r u c t u r e s Report No. 28, Farnborough, England, 1948.

&

11. Garcia, D., Empir ical Formulae f o r Rad i i of Gyra t ion of A i r c r a f t , Revis ion A, SAWE Paper No.78A, Republic Avia t ion Corporat ion, Farmingdale, Long I s l a n d , NY, 1962.

P a r t V References Page 1 2 3

1 2 . Schmit t , R.L., Foreman, K.C., Gertsen, W.M. and Johnson, P.H., Weight Es t imat ion Handbook f o r L igh t Aircraft, Cessna A i r c r a f t Company, 1959.

13. Nicolai , L.M., Fundamentals of ~ i r c r a f t ~ e s i g n , METS, Inc., 6520 Kingsland C o u r t , CA, 95120.

1 4 . Torenbeek, E., Syn thes i s of Subsonic Airplane Design, Kluwer Boston inc . , Hingham, Maine, 19 82.

15. Anon., Fede ra l Aviation Regula t ion , P a r t 23, Depart- ment of T ranspor t a t i on , Fede ra l Avia t ion Administra- t i o n , D i s t r i b u t i o n Requirements Sec t ion , M-482.2, Washington D.C. , 20590.

16. Anon., Federa l Avia t ion Regula t ion , P a r t 25, see R e f . 15.

17. MIL-A-8861(ASG), M i l i t a r y S p e c i f i c a t i o n , Airplane S t r e n g t h and R i g i d i t y , F l i g h t Loads, May 1960.

18. B u s i n e s s and Commercial Av ia t ion (Monthly magazine), 19 85 Planning and Purchasing Handbook, A p r i l 19 85.

1 9 . Roskam, J., Ai rp lane F l i g h t Dynamics and A u t o m a t i c F l i g h t Con t ro l s , P a r t I , Roskam Avia t ion and Engi- neer ing Corporat ion, R t 4 , Box 274, O t t a w a , Kansas, 66067.

P a r t V References Page 124

I X A : DATA S-E FOR COMPONENT TIOm

Tables A 1 through A13 p resen t component weight d a t a f o r t h e following types of a i rp lanes :

1. Homebuilt p rope l l e r dr iven a i r p l a n e s : Tables A l . 1.

2. S ingle engine p rope l l e r dr iven a i rp lanes : Tables A2.1 and A2.2.

3. Twin engine p rope l l e r d r iven airplanes: Tables A3.1 and A3.2.

4. Agr icu l tu ra l a i rp lanes : Tables A4.1. A t t h e time of p r i n t i n g no d a t a were ava i l ab le . The reader should use Tables A2 and add spray equipment weights.

5. B u s i n e s s Jets: Tables AS. 1 and AS. 2.

6. Regional turbopropel le r a i rp lanes : Tables A6.1 through A6.3. Regional p i s t o n / p r o p e l l e r a i rp lanes : Tables A6.4.

7. Je t t r a n s p o r t s : Tables A 7 . 1 through A7.5. Turbopropeller dr iven t r a n s p o r t s : Tables A7.6.

8. Mi l i t a ry t r a i n e r s : Tables A8.1.

9. F ighters : Tables A 9 . 1 through A9.5.

10. M i l i t a r y j e t t r a n s p o r t s : Tables A. 10.1. M i l i t a r y turbopropel le r dr iven t r a n s p o r t s : Tables A10.2. M i l i t a r y p i s t o n / p r o p e l l e r dr iven t r a n s p o r t s : Table A10.3 and A10.4. Mi l i t a ry p a t r o l a i rp lanes : Tables A1O. 5 .

11. Flying boats , amphibious and f l o a t a i rp lanes : Tables A l l . l . A t t h e t i m e of p r i n t i n g no d a t a were ava i l ab le . The reader should use s u i t a b l e t a b l e s i n ca tegor ies 1-10 and account f o r h u l l weight with t h e method of Chapter 4.

1 2 . Supersonic cruise a i rp lanes : Tables A12.1.

13. NACA and NASA X (experimental) a i r p l a n e s : s-

Tables A13.1 through A13.4. CAUTIQN: most of t h e X a i r p l a n e s were b u i l t f o r experimental purposes only. They should n o t be regarded as 'optimized' f o r a given mission.

P a r t V Appendix A Page 125

Table Al . l a Group Weight Data f o r Homebuilt P r o p e l l e r -- -,3=1==I=P=IPPPPIIP~PPPPPPsIP~PPPPPIPOPPI===~======a=

Driven A i r p l anes P = ¶ P I P P P I I I P P P P s

i

TYPe Bede A t t h e time of p r i n t i n g no BD5B o t h e r d a t a were a v a i l a b l e

Weight Item, l b s

Wing Group 8 7 Empennage Group 1 7 Fuselage Group 8 9 Nace l le Group 0 Landing Gear Group 32

Nose Gear 1 0 Main Gear 22 ---

S t r u c t u r e T o t a l 225 --- Engine 1 4 6 A i r Induct . System 0 Fue l System 25 P r o p e l l e r I n s t a l l . 5 Thrus t At tenua tor 3 Engine I n s t a l l . 1 0 --- Power P l a n t T o t a l 189 --- Avionics + Instrum. 1 5 Su r face Con t ro l s 0 E l e c t r i c a l System 1 0 E l e c t r o n i c s 0 B a l l a s t 30 Parachute 20 Furn ish ings + p a i n t 50 Aux i l i a ry Gear 0 --- Fixed Equipm't T o t a l 1 2 5 ---

Fuel 340 Pay load (p i1o t ) 1 7 0

P a r t V Appendix A Page 1 2 6

Table A l . l b Group Weight Data f o r Homebuilt P r o p e l l e r P P P ~ = I P = t l P I P P S f P P I P 5 ~ = I ~ ~ 3 C P ~ P P ~ t ~ ~ P I I C P W = E ~ E ~ ~ ~ ~ ~ ~ ~ ~ 6

Driven Airplanes P=P='PuPPtSPPPPP

Bede A t t h e time of p r i n t i n g no BDSB o t h e r d a t a were a v a i l a b l e

F l i g h t Design Gross Weight, GW, l b s 1 ,051

StructureJGW 0.214 Power Plant/GW 0.180 Fixed Equipmet/GW 0.119 Empty Weight/GW 0.513

Wing GroupIGW 0.083 Empenn. GroupIGW 0.016 Fuse lage GroupIGW 0.0 85 Nacelle Group/GW 0.000 Land. Gear Group/GW 0.030

Take-off Gross Wht, WTOn I b s 1 ,051

Empty Weight, WE, l b s 53 9

Wing GrouplS, p s f 1.8 Emp. GrplSemp, p s f 1.1

U l t i m a t e Load F a c t o r , g ' s 5.7 assumed

S u r f a c e Areas, f t2

Wing, S 47.4 Horiz. T a i l . Sh 10.5

V e r t . T a i l , Sv 5.0

Empenn.

P a r t V

Area, S e v

Appendix A Page 1 2 7

Table A 2 . l a Group Weight Data f o r S i n g l e Engine P r o p e l l e r = = = = = = ~ = = P S I P P P ~ ~ P D = = ~ ~ = ~ = = P P = = : = = = = ~ = = = ~ S ~ = = = = = = = ~ = = = = = =

Driven Ai rp lanes I P I I P P P I I P = P I Z P I

Type ' Cessna 1 5 0 1 7 2 175 180 182 L-19A*

Weight Item, l b s * *

Wing Group Empennage Group Fuselage Group Nace l l e Group Landing Gear Group

Nose Gear Main Gear

S t r u c t u r e To ta l

Engine A i r Induct . System Fue l System P r o p e l l e r I n s t a l l . Engine I n s t a l l .

Power P l a n t T o t a l

Avionics + Instrum. Su r face Cont ro l s E l e c t r i c a l System E l e c t r o n i c s Ai r Cond. System Ant i- icing System Furn i sh ings Aux i l i a ry Gear

Fixed Equipm't T o t a l

Fuel Payload

+ M i l i t a r y o b s e r v a t i o n a i r p l a n e * *Tai ldragger


T a b l e A2. lb Group Weight Data fo r S ing le E n g i n e P r o p e l l e r X I I I I ~ I P I E U D P P P I I P I s m ~ P I I ~ P P E P = P P O = = . L ~ l i P E U E U P = P E ~ = E = = I ~ = =

Driven A i r p l a n e s LIPlPlPPPPIEPPPP

Type C e s s n a 1 5 0 172 175 180 1 8 2 L-19A*

* * F l i g h t Design Gross Weight , GW, l b s 1 , 5 0 0 2 , 2 0 0 2 , 3 5 0 2 , 6 5 0 2 , 6 5 0 2 ,100

S t r u c t u r e I G W 0.406 0.352 0 .330 0.319 0.326 0.327 Power PlantIGW 0.178 0.157 0.177 0.206 0 .206 0.262 F i x e d Equipmet /Gw 0.068 0.072 0.068 0.065 0.065 0 .136 Empty Weight/GW 0.631 0.565 0 .561 0.576 0.583 0.727

Wing Group/GW 0.144 0.103 0.097 0 .089 0.089 0 .113 Empenn. Group/GW 0.024 0.026 0.024 0.023 0.023 0.030 F u s e l a g e GroupIGW 0.154 0.160 0 .149 0.152 0 .151 0.103 Nacelle Group/GW 0.015 0.012 0 .013 0.012 0.013 0.016 Land. Gear GroupfGW 0.069 0 .050 0.047 0 .042 0 .050 0.064

Take- off G r o s s Wht, WTO, I b s 1 , 5 0 0 2 ,200 2 , 3 5 0 2 ,650 2 ,650 2 ,100

Empty Weight, WE, I b 8 946 1 , 2 4 3 1 , 3 1 9 1 , 5 2 6 1 , 5 4 5 1 , 5 2 7

Wing GroupfS , psf 1 . 4 1.4 1 .3 1 .3 1 . 3 1.4 Emp. Grp/Selap, psf 0.85 1.1 1.1 1.2 1 . 2 1 .2

U l t i m a t e Load F a c t o r , g ' s 5.7 5.7 5.7 5.7 5.7 5.7

Surface A r e a s , f t 2

Wing, S 1 6 0 175 175 175 175 174 Horiz. T a i l , Sh 28.5 34.6 34.6 34.6 34.1 35.2

V e r t . T a i l , Sv 14 .1 18.4 18.4 18.4 18.4 18.4

Empenn. Area , S ew 42.6 53.0 53.0 53.0 52.5 53.6

*Military o b s e r v a t i o n airplane ' T a i l d r a g g e r

P a r t V Appendix A P a g e 1 2 9

Table A2.2a Group Weight Data f o r S i n g l e Engine P r o p e l l e r p D I = = P l e P D P S l l l ~ f = I = = = 3 3 1 = 5 = : = = t = = ~ J = = P 5 5 ~ 5 : = ~ = a ~ = = = E ~ = = ~ = = = =

Driven Ai rp lanes 53PSP='==L==I=tP

i

Type Cessna Beech Saab Rockwell Cessna 210A 5-35 S a f i r ll2TCA 210J

Weight I t e m , l b s .

Wing Group Empennage Group Fuselage Group Nace l le Group Landing Gear Group

Nose Gear Main Gear

S t r u c t u r e T o t a l

Engine A i r Induct . System Fuel System P r o p e l l e r I n s t a l l . Engine I n s t a l l .


261 3 79 276 3 34 335 71 5 8 60 9 8 86 316. 200 3 86 358 408. 3 1 62 i n fu s . 61 2 8

207 205 119 161 191 35 50

126 141 .................................. 8 86 904 841 1,082 1,048 .................................. 390 432 47 5 450

3 7 3 0 17 24 73 i n eng. 64 4 5 6 5 3 6 ..................................

577 5 83 5 57 5 81 .................................. Avionics + Instrum. 1 6 16 Sur face Con t ro l s 44 56 i n Hydraul ic System 4 E l e c t r i c a l System 60 72 Air Cond. System 12 1 2 Ant i- ic ing System Furn i sh ings 116 174 Oxygen System 0 0 B a l l a s t 0 0 Auxi l i a ry Gear 0 4 Misc. Equipment 20 0 P a i n t --------------

64 18 fus. 44 4 8

10 51 81 57

i n m i s c . 10

Fixed Equipm't T o t a l 272 334 445 335 ..................................

Fuel (max. payload) 2 34 Payload 84 5

*Inc ludes wing- fuselage carry- through s p a r s *Maximum f u e l


T a b l e A2.2b Group Weight Data for S i n g l e E n g i n e P r o p e l l e r 0 = = X I = P I I I = I S I P P = = = P a 0 t = P e = = ~ Z i P I = P = = = L : = = i = s = = ~ = 0 = = = ~ = = = = = =

Driven A i r p l a n e s =P=P===D====PP=P

C e s s n a Beech S a a b Rockwe l l C e s s n a 210A J-35 S a f i r 112TCA 2 1 0 5

F l i g h t D e s i g n Gross Weight , GW, l b s 2 , 9 0 0 2 , 9 0 0 2 ,660 2 ,954 3 ,400

S t ructure1GW 0.306 0.312 0 .316 0.366 0 .308 Power PlantIGW 0.199 0 .201 0.189 0 .171 F i x e d Equipm't/GW 0.094 0 .115 0 .151 0 .099 Empty Weight /GW 0.598 0 .628 0. 620 0.705 0 .578

Wing GroupIGW 0.090 0 .131 0.104 0 .113 0.099 Empenn. GroupIGW 0.024 0.020 0 .023 0.033 0.025 F u s e l a g e Group/GW 0.109 0.069 0.145 0 .121 0 .120 Nacelle GroupIGW 0.011 0 .021 0 .021 0 .008 Land. Gear GroupIGW 0.071 0 .071 0 .045 0 .055 0.056

Take- of f Gross Wht, WTOs I b s 2 ,900 2 ,900 2 , 6 6 0 2 , 9 5 4 3 ,400

Empty Weight , WE, l b s 1 , 7 3 5 1 , 8 2 1 1 , 6 5 0 2 , 0 8 4 1 , 9 6 4

Wing GKOUP/S, p s f 1 . 5 2 .1 1 . 9 2.2 1.9 Emp. GrplSemp, psf 1 .3 1 .6 1 . 4 2.0 1 .5

U l t i m a t e Load F a c t o r g 's 5. 7

Surface A r e a s , f t2

Wing, S 1 7 6 1 7 8 146 1 5 2 1 7 6 Horiz. T a i l , Sh 38.6 27.6** 32.0 38.6

V e r t . T a i l , S, 17.2 14.3** 17 .0 17.2

Empenn. Area , S emP

55. 8 35.8 41.9 49.0 55. 8


Table A3.la Group Weight Data f o r Twin Engine P r o p e l l e r I X = = P P I I P I P = P ' S ' P S = I = = S = ' ~ = = P = = = X = S ~ ~ P P = = ~ ~ S = = = ~ = ~ ~ ~ ~ ~ P

Driven Ai rp lanes I P I I = P I I D I P I S = I P

'i

TYPe Beech Cessna 65 QA* E-18s G-50 TB* 95 TA* 310C

Number of engines: 2 2 2 2 2 Weight Item, l b s

Wing Group Empennage Group Fuselage Group Nacelle Group Landing Gear Group

Nose Gear Main Gear


Engines A i r Induct . System Fue l System P r o p e l l e r Ins t a l l . Engine Ins t a l l .


Avionics + Instrum. 70 100 80 4 9 46 Sur face Con t ro l s 132 1 1 5 120 73 6 6 Electr ical System 166 295 184 96 121 Electronics 2 63 9 26 0 A i r Cond. System 90 144 81 4 8 46 Ant i- ic ing System Furn ish ings 43 8 524 333 194 154 Aux i l i a ry Gear 5 o 7 0 65 ................................... Fixed Equipm't T o t a l 903 1 ,241 81 4 4 86 49 8 ...................................

Fuel Payload

*QA = Queen A i r , TB = Twin Bonanza, TA Trave l A i r *Ta i ld ragger


T a b l e A3. lb Group Weight Data f o r Twin E n g i n e P r o p e l l e r P I P P I P I P P I S I = P P E I P E P P = O P t = P = P E ~ = . 5 P I = P P Z : E = = E = ~ E = P = ~ ~ = = ~ =

Driven A i r p l a n e s -- --DEIPPPLPPPPPPP

Beech C e s s n a 65 M * E-18S**G-50 TB* 95 TA* 310C

F l i g h t Design Gross Weigh t , GW, l b s 7 ,368 9 , 7 0 0 7 , 1 5 0 4 ,000 4, 830

S t r u c t u r e I G W 0.292 0.282 0.282 0.303 0.265 Power PlantIGW 0.219 0 .235 0 .225 0 .218 0.259 F i x e d Equipmtt/GW 0.123 0 .128 0.114 0.122 0 .103 Empty Weight IGW 0.638 0 .651 0.624 0.649 0 .628

Wing GroupIGW 0.091 0.090 0 .092 0 .115 0.094 Empenn. GroupIGW 0.021 0.019 0 .022 0.020 0.024 F u s e l a g e Group/GW 0.0 82 0.079 0.069 0.069 0.066 Nacelle GroupIGW 0.039 0.034 0.037 0 .045 0.027 Land. G e a r GroupIGW 0.060 0 .060 0 .063 0 .055 0.054

Take- of f Gross Wht, WTO, I b s 7 ,368 9 ,700 7 , 1 5 0 4 ,000 4, 830

Empty Weight, WE, I b s 4 , 7 0 1 6 , 3 1 8 4 , 4 5 9 2 ,595 3 ,032

Wing Group lS , psf 2.4 2.4 2.4 2.4 2.6 Emp. GrplSemp. p s f 1 .4 1 .7 1 .4 1.2 1 . 5

Ultimate Load F a c t o r , g ' s 6.6

S u r f a c e A r e a s , f t 2

Wing, S 27 7 3 6 1 277 194 175 Horiz. T a i l , Sh 79.3 71.6 79.3 42.4 54.3

V e r t . T a i l , Sv 30. 8 33.6 30. 8 23.3 25.9

Empenn. Area , S emp

1 1 0 1 0 5 1 1 0 65.7 80.2

*QA = Queen A i r , TB Twin Bonanza, T A = T r a v e l A i r * T a i l d r a g g e r

P a r t v Appendix A P a g e 1 3 3

Table A3.2a Group Weight Data f o r Twin Engine P r o p e l l e r P====P=OPPPI=P=PI I===P============== I=Bt===========~===

Driven Ai rp lanes ==P=PIZ=====PP== -

TYPe Cessna Rockwell 404-3 414A TP-441 690B

Number of engines: 2 2 2 2 Weight I t e m , l b s (PP) (PP) (PP) (TBP)

Wing Group 86 0 638 87 3 1,001 Empennage Group 181 160 233 207 Fuselage Group 610 678 87 3 1,377 Nacelle Group 2 84 200 258 i n prop/eng Land. Gear Group 316 303 34 6 437

Nose Gear 67 7 5 69 53 Main Gear 249 228 27 7 3 84 ..................................

S t r u c t u r e To ta l 2,251 1,979 2,5 83 3,022 .................................. Engines 1,000 86 2 74 5 720 Air Induct . System 23 3 6 0 17 Fuel System 107 96 9 3 180 P r o p e l l e r I n s t a l l . 215 165 302 Engine I n s t a l l . 2 81 240 130

75 8

.................................. Power P l a n t T o t a l

Avionics + Instrum. Hydraulic System Sur face Con t ro l s Electrical System E l e c t r o n i c s Oxygen System Air Cond. System Ant i- ic ing System Furn i sh ings Aux i l i a ry Gear P a i n t

Fixed Equipm't T o t a l

Fuel Payload***

* Inc ludes p r e s s u r i z a t i o n system **This is a l l b a l l a s t i n t h i s model ***Includes a crew of two


T a b l e A3.2b Group W e i g h t Data for Twin E n g i n e P r o p e l l e r I = P P P = = P I S P I D P P I P P I a O P P ~ = = ~ P i = O ~ L . ~ = = P ~ P = = ~ = = = = ~ P ~ ~ ~ P ~ = =

D r i v e n A i r p l a n e s = S P = = = P P I O = P I P I P

TYPe C e s s n a R o c k w e l l 404-3 414A TP-44 1 690B ( P P ) (PP) (TBP (TBP)

F l i g h t D e s i g n Gross W e i g h t , GW, l b s 8 , 4 0 0 6 , 7 8 5 9 , 9 2 5 1 0 , 2 0 5

S t r u c t u r e / G W 0.268 0 .292 0 .260 0 .296 Power Plant /GW 0.194 0 .206 0 . 1 2 8 0 .164 F i x e d Equipm't/GW 0.134 0 . 1 6 7 0.194 0.187 Empty Weight/GW 0.596 0 .665 0 .582 0 .647

Wing Group/GW 0.102 0.094 0 . 0 8 8 0 . 0 9 8 Empenn. Group/GW 0.022 0.024 0 .023 0 .020 F u s e l a g e Group/GW 0 .073 0 .100 0 .088 0 .135 Nacelle Group/GW 0.034 0 .029 0.026 Land. G e a r Group/GW 0 .038 0 .045 0 .035 0.043

Take- of f Gross Wht, WTO, l b s 8 , 4 0 0 6 , 7 8 5 9 , 9 2 5 1 0 , 2 0 5

Empty W e i g h t ,

W ~ ' l b s 5 , 0 0 6 4 , 5 1 1 5 , 7 8 1 6 , 6 0 5

Wing G r o u p / S , psf 3.6 2. 8 3.4 3. 8 Emp. Grp/Semp. p s i 1 . 7 1 .6 2.2 2.0

U l t i m a t e Load F a c t o r , g ' s 3.75* 3.75* 3.75* 3.75.


Wing, S 242 226 2 5 4 266 Horiz. T a i l , Sh 63.4 60.7 63.4 5 8. 4

V e r t . T a i l , S, 43.5 41.2 43.5 44. 8

Empenn. A r e a , S emp

1 0 7 1 0 2 107 1 0 3


Table A4.1a Group Weight Data f o r Agr icu l tu ra l Airplanes = ~ = P = P I I P ~ I ~ I I D P ~ P I P ~ P P O ~ P I ~ ~ P ~ P O P ~ ~ ~ ~ P P E ~ ~ ~ E ~ ~ E E ~ ~ ~ I U ~ ~

TYPe A t t h e time of p r i n t i n g no d a t a were a v a i l a b l e

~umbe; of engines: Weight Item, l b s

Wing Group Empennage Group Fuselage Group Nacel le Group Landing Gear Group

Nose Gear Main Gear ................................

S t r u c t u r e To ta l ................................ Engines Air Induct. System Fuel System Propulsion System. ................................ Power P lan t To ta l ................................ Spray equipment Avionics + Instrum. Surface Controls Hydraulic System Pneumatic System E l e c t r i c a l System Elec t ron ics Oxygen System A i r Cond. Systeme* Furnishings Auxil iary Gear Miscellaneous ................................ Fixed Equipm't To ta l ................................

Max. F u e l Capacity Max. Payload

P a r t v Appendix A Page 1 3 6

Table A4.lb Group Weight Data f o r Agr icu l tu ra l Airplanes ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ P I P P P I P I I P ~ P P P ~ ~ I ~ I P E ~ P P ~ I ~ U I . L ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ U ~ ~

F l i g h t Design Gross Weight, GW, l b s

S t ruc ture lGW Power Plant/GW Fixed Equipmtt/GW Empty Weight /GW

Wing Group/GW Empenn. Group/GW Fuselage Group/GW Nacelle Group/GW Land. Gear Group/GW

Take-of f Gross Wht, WTO# l b s

Empty Weight,

W ~ ' l b s

Wing Group/S, psf Emp. Grp/S ewe sf

U l t i m a t e Load Factor. g ' s

Surface Areas, f t 2

Wing, S Horiz. T a i l . Sh

V e r t . T a i l . Sv

Empenn. Area, S emp

P a r t V

A t t h e time of p r i n t i n g no d a t a were a v a i l a b l e

Appendix A Page 137

Table A 5 . l a Group Weight Data f o r Business Jets ' P I = = = = = L 3 = 3 P P 3 ~ 5 = P P 5 = = = = = = r D = = = = ~ 3 =

TYPe MS-760 Lockheed Gates- Lear je t S P a r i s Jetstar 2 5 0 2 8

Number of engines: 2 2 2 2 Weight Item, l b s

Wing Group 897 2,827 1,467 1,939 Empennage Group 176 879 361 361 Fuselage Group 912 3,491 1,575 1,624 Nacel le Group 49* 792 241 214 Landing Gear Group 307 1,061 5 84 5 84

Nose Gear 102 102 Main Gear 4 82 4 82 ................................

S t r u c t u r e T o t a l 2,341 9,050 4,228 4,722 ................................ Engines 609 1,750 79 2 792 Air Induct . System 31 135 0 0 Fuel System 240 3 60 179 237 Propuls ion System. 136 230 255 255 ................................ Power P l a n t T o t a l 1,016 2,475 1,226 1,284 ................................ Avionics + Instrum. 70 153 3 83 3 83 Sur face Con t ro l s 18 8 768 2 91 275 Hydraulic System Pneumatic System E l e c t r i c a l System 2 84 973 62 0 603 E l e c t r o n i c s 15 8 86 8 0 0 Oxygen System 2 8 26 A i r Cond. System** 4 8 510 2 93 285 Ant i- ic ing System 8 2 162 Furn ish ings 169 1,521 72 0 768 Auxi l i a ry Gear 0 10 0 0 Miscel laneous 0 0 -4 0 -11 ................................ Fixed Equipm't T o t a l 917 5,065 2,496 2,605 ................................

Max. Fuel Capaci ty 2,460 11,229 6,098 4,684 Max. Payload 8 84 2,100 2,980 1,962

*Engines bu r i ed i n s i d e t h e f u s e l a g e **Includes p r e s s u r i z a t i o n system


T a b l e A5. lb Group Weight Data f o r B u s i n e s s Je ts 'DX l= lP===P I==P%=== IP IPPPDID= I=P==~~aOa======= I~~==~~=~=~=~

MS-760 Lockheed G a t e s - L e a r j e t P a r i s Je ts tar 25D 2 8

F l i g h t Design Gross Weight , GW, l b s 7 ,650 30 ,680 1 5 , 0 0 0 1 5 , 0 0 0

S t r u c t u r e I G W 0.306 0.295 0.282 0.315 Power PlantIGW 0.133 0.081 0.082 0 .086 F i x e d Equipm'tlGW 0.120 0.165 0.166 0.174 Empty Weight IGW* 0.563 0 .541 0.530 0.574

Wing Group/GW 0.117 0.092 0 .098 0.129 Empenn. GroupIGW 0.023 0.029 0.024 0.024 F u s e l a g e GroupIGW 0.119 0.114 0 .105 0.108 Nacelle GroupIGW 0.006* 0.026 0.016 0.014 Land. Gear GroupIGW 0.040 0.035 0.039 0.039

Take- of f Gross W h t ~ WTOs I b s 7 ,650 30 ,680 1 5 , 0 0 0 1 5 , 0 0 0

Empty Weigh t ,

W ~ ' l b s 4 ,306 16 ,590 7 , 9 5 0 8 , 6 1 1

Wing Group lS , psf 4.6 5.4 6.3 7.3 Emp. GrplSemp, psf 3.5 3.4 3.9 3.9

Ultimate Load F a c t o r , g 's 3.75*+ 5.25 3.75** 3.75**

S u r f a c e Areas . f t 2

Wing, S 19 4 521 232 265 B o r i z . T a i l , Sh 31. 8 1 4 9 54.0 54.0

V e r t . T a i l , Sv 18 .4 1 1 0 37.4 37.4


50.2 259 91.4 91.4

* E n g i n e s b u r i e d i n s i d e t h e fuselage **Assumed

P a r t V Appendix A P a g e 139

Table A 5 . 2 a Group Weight Data f o r Business J e t s X J ~ = = D P I = O I P ~ P ~ P P I P = P P = ~ ~ = ~ ~ ~ ~ P I = O ~ = = ~ : P = ~ ~ = ~ ~ = ~

TYPe Cessna Rockwell Hawker- Gul f s t r . 3 C i t a t i o n Siddeley American

11 JC-1121 125 G I I Number of engines: 2 2 2 2 Weight Item, l b s

Wing Group 1,288 1,322 1,968 6,372 Empenn. Group 295 425 60 8 1,965 Fuselage Group 1,069 1,622 1,628 5,944 Nacelle Group 220 350 i n f u s e l . 1,239 Land. Gear Group 465 4 43 65 9 2,011

Nose Gear 87 321 Main Gear 378 1,690 ..................................

S t r u c t u r e Tota l 3,337 4,162 4,863 17,531 .................................. E n g i n e ( ~ ) 1,100 6,570 Air Induct. System 26 Exhaust System 15 Fuel System 189 316 Propulsion System 105 .................................. Power P lan t To ta l 1,435 6,886 .................................. Avionics + Instrum. Surface Controls Hydraulic System E l e c t r i c a l System Elec t ron ics Oxygen Sys tem A i r Cond. System* Anti- icing System Furnishings Auxi l ia ry Gear Auxil iary power u n i t Pa in t

Fixed Equipm't To ta l

Max. Fuel Capacity 5,009 8,964 9,193 23,300 Max. Payload 1,905 5,380

*Includes p r e s s u r i z a t i o n system


T a b l e A5.2b Group Weight Data f o r B u s i n e s s Jets a P P P I S ~ P I a L P I P I I a P l = = = r ~ t a r P P P P P ~ t ~ ~ P ~ t P P = a ~ ~ ~ E

C e s s n a Rockwe l l Hawker ' G u l f s t r . C i t a t i o n S i d d e l e y Amer ican 11 JC-1121 1 2 5 G I I

F l i g h t D e s i g n Gross Weight , GW, l b s 1 3 , 5 0 0 20 ,500 2 3 , 3 0 0 64 ,800

S t r u c t u r e I G W 0.247 0 .203 0.209 0 .271 Power PlantIGW 0.106 0.106 F i x e d Equiprnet/GW 0.167 0.173 Empty Weight lGW 0.520 0.540 0.526 0.550*

Wing GroupIGW 0.095 0.064 0 .084 0 .098 Empenn. GroupIGW 0.022 0.021 0.026 0.030 F u s e l a g e GroupIGW 0.079 0.079 0.070 0.092 Nacelle GroupIGW 0.016 0.017 i n f u s e l . 0.019 Land. Gear GroupIGW 0.034 0 .022 0.028 0 .031

Take- of f Gross Wht. WTO, I b s 1 3 , 5 0 0 20 ,500 23 ,300 64 ,800

Empty Weight , WE. I b s 7 , 0 2 3 1 1 , 0 7 0 1 2 , 2 6 0 35 ,620

Wing GroupIS , psf 4.6 4. 4 Emp. GrplSemp, psf 2.4 3.3

Ultimate Load F a c t o r , g ' s 3.75..

S u r f a c e A r e a s . f t 2

Wing, S 279 303 353 794 Horiz. T a i l . Sh 70.6 70 1 0 0 1 8 2

V e r t . T a i l , S, 50.9 59.3 51.6 1 5 5


1 2 2 1 2 9 1 5 2 337

* T y p i c a l . I n d i v i d u a l airplanes w i l l vary. *Assumed


Table A6.la Group Weight Data f o r Regional Turboprope l le r ==P==IPPPIPP=DPDIItPPLie:==P3E=P=ZiIP=I=L:L=======a======sD===

Driven Ai rp l anes = I P I I P D = = I P = O a I P -

TYPe Grumman Fokker Nord Embraer G- I F-27-100 262 110-P2


Wing Group 3,735 4,408 2,698 1,502 Empennage Group 874 97 7 80 5 454 Fuselage Group 3,718 4,122 3,675 1,354 Nacel le Group 1,136 62 8 236 19 8 Land. Gear Group 1,207 1,940 1,085 538

Nose Gear 219 Main Gear 9 88 ...................................

S t r u c t u r e T o t a l 10,670 12,075 8,499 4,046 ................................... Engines 2,688 2,427 62 2 Air Induct . System Fuel System 133 39 0 8 6 P r o p e l l e r I n s t a l l . 1,002 91 8 1,140 Propuls ion System 69 8 612 i n prop. ................................... Power P l a n t T o t a l 4,521 4,347 1,848 ................................... Avionics + Instrum. 97 81 133 3 64 Sur face Con t ro l s 461 61 3 408 342 Hydraulic System 235 242 ( i n c l . i n 176 Pneumatic System e l e c t r . E l e c t r i c a l System 966 83 5 7 65 452 E l e c t r o n i c s 99 3 86 238 i n avion. APU 355 0 0 0 Air Cond. System Ant i- ic ing System - - Furn ish ings 415 2,291 1,324 8 82 Auxi l i a ry Gear 6 3 3

Fixed Equipm't T o t a l 3,389 5,673 3,428 2,481 ...................................

Max. Fuel Capaci ty 10,447 9,198 3,559 3,062 Maximum Payload 4,270 12,500 6,175 3,706

*Inc ludes p r e s s u r i z a t i o n system

p a r t v Appendix A Page 142

T a b l e A6. lb Group Weight Data f o r R e g i o n a l T u r b o p r o p e l l e r

Driven A i r p l a n e s P P = = = a P I = P I P I P P I

Grumman F o k k e r Nord Embraer G - I F-27-100 262 110-P2

F l i g h t D e s i g n Gross Weigh t , GW, l b s 35 ,100 37 ,500 2 2 , 9 3 0 1 2 , 5 0 0

S t r u c t u r e I G W 0.304 0.322 0 .371 0.324 Power PlantIGW 0.129 0.116 0 .148 F i x e d Equipmtt/GW 0.097 0 .151 0.149 0.19 8 Empty WeightIGW 0.624 0 .615 0.663 0.670

Wing GroupIGW 0.106 0 .118 0 .118 0.120 Empenn. Group/GW 0.025 0.026 0.035 0.036 F u s e l a g e GrouplGW 0.106 0 .110 0 .160 0 .108 Nacelle GroupIGW 0.032 0.017 0 .010 0 .016 Land. Gear GroupIGW 0.034 0.052 0.047 0.043

Take- of f Gross Wht, WTO, I b s 35 ,100 37 ,500 2 2 , 9 3 0 1 2 , 5 0 0

Empty Weight ,

W ~ ' l b s 21 ,900 23 ,054 1 5 , 2 0 0 8 ,3 75

Wing GroupIS , p s f 6.1 5. 8 4. 6 4. 8 Emp. GrplSemp, p s f 3.6 3.0 2.9 2. 8

Ultimate Load F a c t o r , g ' s 3.75* 3.75* 3.75* 3.75*


Wing, S 610 754 592 313 Horiz. T a i l , Sh 127 172 169 1 0 5

V e r t . T a i l , Sv 117 1 5 3 109 5 9


244 325 27 8 164

P a r t v Appendix A P a g e 143

Table A6.2a Group Weight Data f o r Regional Turbopropeller ~ = ~ ~ P ~ P P I I P = ~ P I X I P I ~ P P P ~ P P P ~ ~ ~ ~ P D ~ P P ~ P P ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ P ~ ~ ~ ~ P ~ ~

Driven Airplanes PPPPPPIIIPPPPI=P

* TYPe Fokker Short*

F-27-200 F-27-500 Skyvan Number of engines: 2 2 2 Weight Item, l b s

Wing Group Empennage Group Fuselage Group Nacelle Group Land. Gear Group

Nose Gear Main Gear

S t r u c t u r e Tota l

Engines Air Induct. System Fuel System Prope l l e r I n s t a l l . Propulsion System

Power P l a n t Tota l

Avionics + Instrum. Surface Controls Hydraulic System Pneumatic System E l e c t r i c a l System Elec t ron ics APU Air Cond. System Anti- icing System Furnishings Auxi l ia ry Gear P a i n t

Fixed Equipm't To ta l

Max. Fuel Capacity 9 , 1 4 6 Payload (Max. 1 2 , 6 1 5

4 , 9 2 4 1 2 , 3 83

*Strutbraced wing **Includes p r e s s u r i z a t i o n ***Cockpit fu rn i sh ings only


T a b l e A6.2b Group Weight Data for R e g i o n a l T u r b o p r o p e l l e r U P I ~ P P P I P ~ P I I I = P P P I I . C . ~ P P ~ I I L ~ ~ ~ E ~ ~ P P P P ~ ~ ~ ~ ~ ~ I = ~ ~ ~ E ~ = ~ ~ P ~ ~

Driven A i r p l a n e s P 5 P P ~ a P I D I P I P I P P

F o k k e r S h o r t * F-27-200 F-27-500 Skyvan

F l i g h t Design Gross Weigh t , GW, l b s 43 ,500 4 5 , 0 0 0 1 2 , 5 0 0

S t r u c t u r e / G W 0.284 0.294 0.357 Power Plant /GW 0.123 F i x e d EquipmVt/GW 0.144 0.086 Empty Weight /GW 0.537 0 .548 0.570

Wing GrouplGW 0.104 0.100 0 .098 Empenn. Group/GW 0.024 0.024 0.030 F u s e l a g e Group/GW 0.099 0.114 0.172 Nacelle Group/GW 0.015 0.015 0.020 Land. G e a r GrouplGW 0.042 0 .041 0 .037

Take- of f Gross Wht, WTOD I b s 43 ,500 45 ,000 1 2 , 5 0 0

Empty Weight, WE, l b s 23 ,350 24 ,650 7 ,125

Wing Group/S , psf 6.0 6.0 3.3 Emp. Grp/SempD psf 3.2 3.3 2.2

Ultimate Load F a c t o r , g ' s 3.75** 3.75** 3.75**


Wing, S 754 75 4 373 H O K ~ Z . T a i l . Sh 172 172 8 5

V e r t . T a i l , Sv 1 5 3 1 5 3 83


325 325 1 6 8

* S t r u t b r a c e d wing **Assumed


Table A 6 . 3 a Group Weight Data f o r Regional Turboprope l le r = = = I P = P P = = P S P P = P P P P = ~ = f = I a = = P Z i C = I a = = = = = l i = = P = = = = = = = = = = = = = =

Driven Ai rp lanes ------a - ------ -==I=====

4

TYPe D e Havil land Canada DHC7-102 DHC6-300

Number of engines: 2 2 Weight Item, l b s

Wing Group 4,888 1,263. Empennage Group 1,318 303 Fuselage Group 4,680 1,705 Nacel le Group 1,841 221 Land. Gear Group 1,732 613

Nose Gear Main Gear ----------------

S t r u c t u r e T o t a l 14,459 4,105 ---------------- Engines A i r Induct . System Fuel System P r o p e l l e r I n s t a l l . P ropuls ion System ---------------- Power P l a n t T o t a l 4,701 1,248 ---------------- Avionics + Instrum. 85 0 371 Sur face Con t ro l s 710 145 Hydraulic System 493

43 Pneumatic System E l e c t r i c a l System 1,651 356 E l e c t r o n i c s A i r Cond. System 550

103** P r e s s u r i z a t i o n System Ant i- ic ing System 176 Furn ish ings 2, 862 732 P a i n t 150 64 ---------------- Fixed Equipm't T o t a l 7,442 1,814 ----------------

F u l l o i l 130 5 4 Max. Fue l Capaci ty 6,968 1,114 Water and s u p p l i e s 130 Payload (Max.) 9,500 3,610

*St ru tbraced wing **Heating system on ly


T a b l e A6.3b G r o u p W e i g h t Data f o r R e g i o n a l T u r b o p r o p e l l e r = = = P P I = P P P = = P P P = P P = = a a = P = = I P S = = = = = = P P = = = = = = = = =

Driven A i r p l a n e s = I I = P P n I P r P a t = = =

De H a v i l l a n d C a n a d a DHC7-102 DHC6-300

F l i g h t D e s i g n G r o s s W e i g h t , GW, l b s 4 4 , 0 0 0 1 2 , 5 0 0

S t r u c t u r e I G W 0.329 0 .328 Power Plant /GW 0.107 0 .100 F i x e d Equipm'tIGW 0.169 0 .145 Empty ~ e i g h t 1 G W 0.605 0 .573

Wing GroupIGW 0 .111 0.101* Empenn. GroupIGW 0.030 0.024 F u s e l a g e G r o u p l G ~ 0.106 0 .136 Nacelle Group/GW 0.042 0 . 0 1 8 Land. Gear GroupIGW 0.039 0 .049**

T a k e- o f f G r o s s Wht, WTO, I b s 4 4 , 0 0 0 1 2 , 5 0 0

Empty Weight, WE, l b s 2 6 , 6 0 2 7 , 1 6 7

Wing G r o u p / S , psf 5.7 3.0 Emp. Grp/Semp, psf 3 .4 2 .1

Ultimate Load F a c t o r , g t s 3.75*** 3.75***


Wing, S 86 0 420 Horiz. T a i l , Sh 217 1 0 0

V e r t . T a i l , SV 1 7 0 4 8


* S t r u t b r a c e d wing * * F i x e d g e a r *Assumed


Table A6.4a Group Weight Data f o r Regional P is ton1 = = P I I P = P t l P I P P P I P P E E P P ~ ~ O P P P P . L P 3 5 = = P f P = P = = = = ~ = ~ = P = ~

Prope l l e r Driven Airplanes a I I 6 P P I I P P O I I P I P t P O P ~ ~ P ~ a E

Handley Sco t t i sh* Type' SAAB Page Aviation Convai r

Scandia Herald Twin Pion. 240 Number of engines: 2 4 2 2 Weight Item, l b s

Wing Group 4,195 4,365 2,121 3,943 Empennage Group 5 84 9 87 57 6 922 Fuselage Group 2,773 2,986 1,381 4,227 Nacelle Group 1,479 83 0 230 1,215 Land. Gear Group 1,841 1,625 703 1,530

Nose Gear Main Gear ..................................

S t r u c t u r e To ta l 10,872 10,793 5,011 11,837 .................................. Engines Air Induct. System F u e l System Prope l l e r I n s t a l l . Propulsion System .................................. Power P lan t To ta l 7,299 .................................. Avionics + Instrum. Surface Controls 3 69 3 64 300 Hydraulic System Pneumatic System E l e c t r i c a l System Elec t ron ics APU Oxygen System A i r Cond. System Anti- icing System Furnishings Auxil iary Gear .................................. Fixed Equipm't To ta l 4,444 ..................................

not-------------------------- 'oil+ 'ti0

known

Max. Fuel Capacity 11, 080 1,740 6,700 Payload (Xax. 1 2,950 16,000

*Strutbraced wing

Part V Appendix A Page 148

T a b l e A6.4b Group Weigh t Data fo r R e g i o n a l P i s t o n 1 I ~ P P L I a P I P I ~ I I P = P = I I = I : I P I P I I P t O I I = ~ I I I I P = = = ~ I ~ = = ~ E

P r o p e l l e r Driven A i r p l a n e s E t l P P I ~ P I I P P I U I I I P I I P P I I L P P

Handley S c o t t i s h TYPe SAAB P a g e A v i a t i o n Conva i r

S c a n d i a H e r a l d Twin P ion . 240

F l i g h t Design Gross Weigh t , GW, l b s 30 ,860 37 ,500 1 4 , 6 0 0 4 3 , 5 0 0

S t r u c t u r e I G W 0.352 0 .288 0.343 0.272 Power PlantIGW 0 .168 F i x e d Equipm'tIGW 0.102 Empty WeightIGW 0.641 0.673 0.683 0.542

Wing Group/GW 0.136 0.116 0.145 0.091 Empenn. GroupIGW 0.019 0.026 0.039 0 .021 F u s e l a g e GroupIGW 0.090 0.080 0.095 0.097 Nacelle GroupIGW 0.048 0.022 0.016 0 .028 Land. Gear GroupIGW 0.060 0.043 0 .048 0.035

Take- of f Gross Wht, WTO, I b s 30 ,860 37 ,500 1 4 , 6 0 0 43 ,500

Empty Weight, WE, l b s 19 ,780 25 ,240 9 ,969 2 3 , s 80

Wing GroupIS, p s f 4.5 4.9 3.2 4. 8 Emp. G r p / S e m p ~ p8f 2.0 2.2 1.7

U l t i m a t e Load F a c t o r , g ' s 3.75* 3.75. 3.75. 3.75*


Wing, S 922 8 86 6 70 81 7 Horiz. T a i l , Sh 215 252 1 6 7

V e r t . T a i l , Sv 82 1 9 3 1 6 7


2 97 445 3 34

*Assumed


Table A 7 . l a Group Weight Data f o r Jet Transpor t s --=---- ,, - - - - P P P P P P = I P = I P P P I = = P ~ n P P = I P = = = i = % ~ P P = = P n = = = =

Type McDonnell Douglas -i DC-9-30 MD-80 DC-10-10 DC-10-30

Number of engines: 2 2 3 3 Weight I t e m , l b s

Wing Group 1 1 , 4 0 0 1 5 , 5 6 0 4 8 , 9 9 0 5 8, 859 Empennage Group 2 , 7 8 0 3 , 3 2 0 1 3 , 6 6 0 1 4 , 6 7 6 Fuselage Group 1 1 , 1 6 0 1 6 , 1 5 0 4 4 , 7 9 0 4 7 , 2 7 0 Nacelle Group 1 , 4 3 0 2 , 1 2 0 8 , 4 9 0 9 , 1 2 7 Land. Gear Group 4 , 1 7 0 5 , 3 4 0 1 9 , 8 2 0 2 5 , 7 6 1

Nose Gear 470 5 5 0 1 , 5 2 0 1 , 8 3 2 Main Gear 3 , 7 0 0 4 , 7 9 0 1 8 , 3 0 0 2 3 , 9 2 9 .

S t r u c t u r e T o t a l 3 0 , 9 4 0 4 2 , 4 9 0 1 3 5 , 7 5 0 1 5 5 , 6 9 3 .................................. E n g i n e ( ~ ) 6 , 4 1 0 8, 820 2 3 , 6 8 8 2 6 , 1 6 3 Exhaust and Thrus t Reverser System 1 , 2 4 0 1 , 5 4 0 7 , 2 3 2 6 , 9 1 6 A i r Induct . System 0 0 0 0 Fuel System 6 0 0 64 0 2 , 0 4 0 4 , 3 0 8 Propuls ion I n s t a l l . 0 0 0 0


Avionics + Instrum. Su r face Cont ro l s Hydraulic System Pneumatic System E l e c t r i c a l System E l e c t r o n i c s APU Oxygen System Air Cond. System** Ant i- ic ing System Furn ish ings Operat ing Items

8 , 2 5 0 1 1 , 0 0 0 3 2 , 9 6 0 3 7 , 3 8 7 .................................. 1 , 4 5 0 2 , 1 3 0 3 , 4 1 0 4 , 2 7 4 1 , 6 2 0 2 , 5 4 0 5 , 8 8 0 6 , 0 1 0

4 80 5 4 0 2 , 3 3 0 2 , 5 8 7 2 80 2 9 0 1 , 7 9 0 1 , 9 2 0

1 , 3 3 0 1 , 7 2 0 5 , 3 7 0 5 , 9 1 2 Included i n Avionics and Instrum.

82 0 84 0 1 , 5 9 0 1 , 6 4 3 1 5 0 2 2 0 2 1 0 2 56

1 , 1 2 0 1 , 5 8 0 2 , 3 9 0 2 , 7 2 3 4 80 5 5 0 4 2 0 47 1

8 , 4 5 0 1 1 , 4 0 0 3 5 , 8 1 0 3 4 , 1 2 4 2 , 7 0 0 3 , 6 5 0 1 3 , 3 4 0 1 6 , 2 7 4

Fixed Equipm't T o t a l 1 8 , 8 8 0 2 5 , 4 6 0 . 7 2 , 5 4 0 7 6 , 1 9 4 .................................. W t f o

Not ......................... known

Max. Fuel Capaci ty 2 8 , 7 4 6 3 9 , 3 6 2 1 4 6 , 6 8 3 2 4 7 , 0 3 4 Max. Payload 2 8 , 9 3 0 4 3 , 0 5 0 9 3 , 7 5 0 9 8 , 7 2 6

*Inc ludes 3 , 5 9 0 l b s f o r c e n t e r l i n e g e a r **Inc ludes p r e s s u r i z a t i o n system


T a b l e A7.lb Group Weight Data f o r J e t T r a n s p o r t s = P P I P P I I I P t P D P D l t ~ 5 O ~ ~ P ~ P ~ P P P E . 5 ~ % P P P ~ ~ ~ ~ ~ ~ ~ ~ = ~ ~

Mc Donne11 D o u g l a s DC-9-3 0 MD-80 DC-10-10 DC-10-3 0

F l i g h t D e s i g n Gross Weigh t , GW, l b s 108 ,000 1 4 0 , 0 0 0 430 ,000 555 ,000

S t ructure/GW 0.286 0.304 0.316 0.281 Power Plant lGW 0.076 0.079 0.077 0.067 F i x e d Equipmet/GW 0.175 0.182 0.169 0.137 Empty WeightIGW 0.538 0.564 0 .561 0.485

Wing GroupIGW 0.106 0 .111 0.114 0.106 Empenn. Group/GW 0.026 0.024 0.032 0.026 F u s e l a g e GrouplGW 0.103 0.115 0.104 0.085 Nacelle GroupIGW 0.013 0.015 0.020 0.016 Land. Gear GroupIGW 0.039 0.038 0.046 0.046

Take- of f Gross Whtr WTOn I b s 1 0 8 , 0 0 0 14OD0O0 430 ,000 555 ,000

Empty Weight, WED l b s 58,070 78 ,950 241 ,250 269,274

Wing GroupIS , psf 11 .4 12.3 12.7 14.9 Emp. GrplSemp, p s f 6.4 5.7 7.0 7.6

Ultimate Load F a c t o r , g 's 3.75* 3.75* 3.75. 3.75*


Wing, S 1 , 0 0 1 1 , 2 7 0 3 , 8 6 1 3 , 9 5 8 Horiz. T a i l , Sh 27 6 3 1 4 1 , 3 3 8 1 , 3 3 8

V e r t . T a i l , S, 1 6 1 1 6 8 605 605

Empenn. Area, S emp

43 7 5 82 1 , 9 4 3 1 , 9 4 3


Table A7.2a Group Weight Data f o r J e t Transports ~ P P O I I I P I P I E I P P a P P I P % t P P P P O P O P I I P P 3 . P P P I P ~ ~ ~ n ~ ~ ~ a

TYPe Boeing Airbus i 737-200 727-100 747-100 A-300 B2


Wing Group 10,613 17,764 86,402 44,131 Empennage Group 2,718 4,133 11,850 5,941 Fuselage Group 12,108 17,681 71,845 35,820 Nacelle Group 1,392 3,870 10,031 7,039 Land. Gear Group 4,354 7,211 31,427 13,611


S t r u c t u r e To ta l 31,185 50,659 211,555 106,542

Engines 6,217 9,325 34,120 16,825 Exhaust and Thrust- Reverser System 1,007 1,744 6,452 4,001 A i r Induct. System 0 0 0 0 Fuel System 575 1,143 2,322 1,257 Propulsion I n s t a l l . 378 250 80 2 81 4 .................................. Power P lan t To ta l 8,177 12,462 43,696 22,897 .................................. Avionics + Instrum. 62 5 756 1,909 377 Surface Controls 2,348 2,996 6,9 82 5, 808 Hydraulic System Pneumatic System E l e c t r i c a l system 1,066 2,142 3,348 4,923 Elec t ron ics 956 1,591 4,429 1,726 APU 83 6 60 1,130 9 83 A i r Cond. System* 1,416 1,976 3,969 3,642 Anti- ic ins System - - Furnishings 6,643 10,257 37,245 13,161 Miscellanous 124 8 5 -421 732 .................................. Fixed Equipm't To ta l 14,887 21,281 63,062 35,053 ..................................

'oil+ W t f o Not ......................... known

Max. Fuel Capacity 34,718 48,353 331,675 76,512 Max. Payload 34,790 29,700 140,000 69,865



T a b l e A7.2b Group Weight Data for Je t T r a n s p o r t s P I D P P S P I I I I P P m ~ P ~ I ~ P P = P = P P ~ I P ~ I I E I t . c I P ~ P ~ m = ~ = ~ = ~

Boeing A i r b u s 737-200 727-100 747-100 A300-B2

F l i g h t Design Gross Weight , GW, l b s 1 1 5 , 5 0 0 160 ,000 710 ,000 302 ,000

S t r u c t u r e I G W 0.270 0.317 0 .298 0.353 Power PlantIGW 0.071 0 .078 0.062 0.076 F i x e d Equipmtt /Gw 0.129 0.133 0.089 0.116 Empty Weight /GW 0.521 0.552 0 .498 0.559

Wing GroupIGW 0.092 0 .111 0 .122 0.146 Empenn. GroupIGW 0.024 0.026 0.017 0.020 F u s e l a g e Group/GW 0.105 0 .111 0 .101 0.119 Nacelle GroupIGW 0.012 0.024 0.014 0.023 Land. Gear G r o u p / G ~ 0 .038 0 .045 0.044 0 .045

Take- off Gross Wht, WTO, l b s 115 ,500 160 ,000 710 ,000 302 ,000

Empty Weight, W E D l b s 60 ,210 88,300 353 ,398 1 6 8 , 8 0 5

Wing GroupIS , psf 10 .8 10.4 15.7 15. 8 Emp. Grp/S .lop# psf 4.9 5.6 5.2 4. 8

ultimate Load F a c t o r , g ' s 3-75. 3.758 3.75. 3.75.


Wing, S 9 80 1 , 7 0 0 5 ,500 2 ,799 Horiz. T a i l . Sh 3 2 1 376 1 , 4 7 0 74 8

V e r t . T a i l , S, 233 356 83 0 4 87


554 732 2 ,300 1 , 2 3 5


Table A 7 . 3 a Group Weight Data f o r Jet Transpor t s -=------=ppp=- - ------ -P=P=PPP==IPPPPPPPPtPP=PIP=P==a=re:P=====r==

Type Boeing -?s 707-121 707-320 707-320C 720-022

Number of eng ines : 4 4 4 4 Weight Item, l b s

Wing Group 24,024 29,762 32,255 22,850 Empennage Group 5,151 5,511 6,165 5,230 Fuselage Group 20,061 21,650 26,937 19,035 Nacel le Group 4,639 4,497 4,183 4,510 Land. Gear Group 9,763 12,700 12,737 8,110


S t r u c t u r e T o t a l 63,638 74,120 82,277 59,735 .................................. Engines 16,458 20,200 17,368 13,770 Exhaust and Thrust - Reverser System 3,492 Air Induct . System 0 0 0 0 Fuel System 1, 808 2,418 1,240 Propuls ion I n s t a l l . 1,738 79 8 8 85 .................................. Power P l a n t T o t a l 20,004 24,076 15,895 .................................. Avionics + Instrum. 505 515 5 5 5 Sur face Cont ro l s 2,159 2,400 3,052 2,450 Hydraulic System Pneumatic System 4 84

E l e c t r i c a l System 3,772 E l e c t r o n i c s 1,708 APU A i r Cond. System* 3,110 Anti- icing System Furn ish ings 13,651 ~uxiliary- ear 0 0 0 0 Miscel lanous 0 0 -3 89 0 .................................. Fixed Equipm't T o t a l 25,389 24,456 24,725 ..................................

Max. Fuel Capaci ty 90,842 160,783 160,783 99,954 Max. Payload 42,600 55,000 84,000 2 8,200



T a b l e A7.3b Group Weight Data f o r Je t T r a n s p o r t s = = = = = = I P P I x P P L ~ I P E ~ L P P P P L P ~ ~ ~ ~ ~ ~ ~ = E ~ ~ I P I ~ ~ ~ ~ ~ ~ ~ ~

Boeing 707-121 707-320 707-32OC 720-022

F l i g h t D e s i g n Gross Weigh t , GW, l b s 246 ,000 311 ,000 330 ,000 203 ,000

S t r u c t u r e I G W 0.259 0 .238 0 .249 0.294 Power PlantIGW 0.081 0 .073 0 .078 F i x e d Equipm'tIGW 0.103 0.074 0.122 Empty Weight/GW 0.444 0.434 0.396 0.494

Wing GroupIGW 0 .098 0.096 0 .098 0.113 Empenn. GroupIGW 0.021 0 .01 8 0.019 0.026 F u s e l a g e GroupIGW 0.0 82 0.070 0.082 0.094 Nacelle GroupIGW 0.019 0.014 0.013 0.022 Land. Gear GroupIGW 0.040 0 .041 0.039 0 .040

Take- of f Gross Wht, WTO, l b s 246 ,000 3 1 1 , 0 0 0 330 ,000 203 ,000

Empty Weigh t , WE, l b s 1 0 9 , 1 1 1 1 3 5 , 0 0 0 130 ,809 1 0 0 , 3 5 5

Wing GroupIS , psf 9.9 10 .3 10.6 9 .4 Emp. GrplSernp, p s f 6.2 5. 8 6.5 6.3

Ultimate Load F a c t o r , g ' s 3.75 3.75 3.75 3.75


Wing, S 2 ,433 2 ,892 3 ,050 2 ,433 Horiz. T a i l , Sh 500 62 5 625 500

V e r t . T a i l , S, 328 328 328 328


82 8 953 9 53 82 8


Table A7.4a Group Weight Data f o r Jet Transports = X I I a I I I P P u u P P I P P I P n P P P L : P I m I ~ t E u P P ~ P 3 . . I u ~ P u ~ ~ a ~ ~ ~

Hawker- TYPe Boeing McDonnell Douglas Siddeley - 707-321 DC-8 DC-9-10 121-IC Number of engines: 4 4 2 3 Weight I t e m , l b s

Wing Group 28,647 27,556 9,470 12,600 Empennage Group 6,004 4,840 2,630 3,225 Fuselage Group 22,129 19,910 11,206 12,469 Nacel le Group 5,119 3,534 1,417 i n f u s e l . Land. Gear Group 11,122 10,910 3,660 4,413


S t r u c t u r e Tota l 73,021 66,750 28,383 32,707 .................................. E n g i n e ( ~ ) 19,192 6,160 Exhaust and Thrust Reverser System 65 8 A i r Induct. System Fuel System 1,956 510 Propulsive I n s t a l l . 1,113 409 .................................. Power P l a n t To ta l 22,261 27,677 7,737 .................................. Avionics + Instrum. 561 719 Surface Controls 2,408 1,264 1,792 Hydraulic System Pneumatic System 498 714

E l e c t r i c a l System 3,959 1,663 Elec t ron ics 1,716 914 APU 81 8 Air Cond. System* 3,290 Anti- icing System Furnishings 14,854 Auxil iary Gear 0 24

Fixed Equipm't To ta l 27,286 25,650 15,000 ..................................

Max. Fuel Capacity Max. Payload



T a b l e A7.4b Group Weight Data for Je t T r a n s p o r t s P I S P P I P I P ~ ~ I I I I E L I n n P . L P P I ~ P P ~ ~ P P E P P ~ = P P = = E = ~ = = ~ E

Hawker Boe ing McDonnell D o u g l a s S i d d e l e y 707-321 DC-8 DC-9-10 121- IC

F l i g h t Design G r o s s We igh t , GW. l b s 302 ,000 215 ,000 91 ,500 1 1 5 , 0 0 0

S t r u c t u r e I G W 0.242 0.310 0.310 0.2 84 Power PlantIGW 0.074 0.129 0.085 F i x e d Equipmet/GW 0.090 0.119 0.164 Empty WeightIGW 0.406 0.562 0.495 0 .587

Wing GroupIGW 0.095 0 .128 0.103 0.110 Empenn. GroupIGW 0.020 0 .023 0.029 0 .028 F u s e l a g e G r o u p I G ~ 0.073 0.093 0 .122 0 .108 Nacelle GroupIGW 0.017 0.016 0.015 i n fusel . Land. Gear GroupIGW 0.037 0 .051 0.040 0.038

Take- of f Gross Wht. WTOn I b s 301 ,000 215 ,000 91 ,500 1 1 5 , 0 0 0

Empty Weight, W E D l b s 122 ,509 1 2 0 , 8 7 7 4 5 , 3 0 0 67 ,500

Wing GroupIS , psf 9.9 9.9 10. 1 9.3 E q . Grp/Semp, psf 6.4 5.6 5. 5 5.7

U l t i m a t e Load F a c t o r , g ' s 3.75 3.75' 3.758 3.758


Wing, S 2 ,892 2 , 7 7 3 93 4 1 , 3 5 8 Horiz. T a i l . Sh 62 5 607.8 275 310

V e r t . T a i l , S, 312 263.8 200.8 259


937 87 0 47 5 569

*Assumed * * E s t i m a t e d from threeview


Table A 7 . 5 a Group Weight Data f o r Je t Transpor t s --- ~ ~ ~ ~ = = ~ P ~ P P I I I P P P I I ~ I ~ O ~ ~ P P P ~ ~ O ~ ~ ~ I ~ : P ~ ~ ~ ~ ~ U ~ E ~ ~ ~

VFW- SudlAe ro- TYPe Fokker Fokker BAC s p a t i a l e

614 F2 8- 1000 1- 111300 Carave l l e i umber? of eng ines : 2 2 2 2 Weight Item, l b s

Wing Group Empennage Group Fuselage Group Nace l le Group Land. Gear Group

Nose Gear Main Gear


Engines Exhaust and Thrust- Reverser System Air Induct . System Fuel System Propu l s ive I n s t a l l


Avionics + Instrum. Su r face Con t ro l s Hydraulic System Pneumatic System E l e c t r i c a l System E l e c t r o n i c s APU

5 , 7 6 7 7 , 3 3 0 9 , 6 4 3 1 4 , 7 3 5 1 , 1 2 1 1 , 6 3 2 2 , 3 6 9 1 , 9 5 7 5 , 2 3 3 7 , 0 4 3 9 , 7 1 3 1 1 , 5 7 0

9 7 1 834 i n f u s e l . 1 , 5 8 1 1 , 6 2 0 2 , 7 5 9 2 , 8 5 6 5 , 1 1 0

Air Cond. System* 719 1 , 0 7 4 1 , 5 7 9 1 , 7 5 2 Anti- icing System Furn ish ings 2 , 6 5 5 4 , 0 3 0 4 , 9 3 3 6,4 81 Auxi l i a ry Gear 4 9 0 0 0 o p e r a t i n g I tems ................................... Fixed Equipm't T o t a l 6 , 5 8 1 9 , 3 9 5 1 2 , 9 5 1 1 5 , 9 4 1 ................................... W t f o

not------------------------- known

Max. Fuel Capaci ty 1 0 , 1 4 2 1 7 , 3 3 1 2 4 , 9 5 4 33 , 808 Max. Payload 8 , 2 0 1 1 4 , 3 8 0 2 2 , 2 7 8 2 9 , 1 0 0



T a b l e A7.5b Group Weight Data f o r Je t T r a n s p o r t s = P I D P = = t P I P = t I P t e P P = = P = = = P = P ~ = a = = = P = I a I = ~ = = = = = ~ =

VFW Sud-Aero TYPe Fokke r F o k k e r BAC spat iale

614 F28-1000 1-111300 C a r a v e l l e

F l i g h t D e s i g n Gross Weight , GW, l b s 4 0 , 9 8 1 65 ,000 87,000 110 ,230

S t r u c t u r e / G W 0.359 0.302 0.283 0.317 Power PlantIGW 0.107 0 .083 0.079 F i x e d Equipm*t/GW 0.161 0 .145 0.149 0.145 Empty Weight /GW* 0.586 0.480 0.560 0.590

Wing GroupIGW 0.141 0 .113 0.111 0.134 Empenn. GrouplGW 0.027 0.025 0.027 0 .018 F u s e l a g e GroupIGW 0.128 0 .108 0.112 0.105 Nacelle Group/GW 0.024 0.013 i n fusel . 0.014 Land. Gear GroupIGW 0.040 0.042 0 .033 0.046

Take- of f Gross Wht, WTOs l b s 40 ,981 65 ,000 87 ,000 110 ,230

Empty Weight , WE, I b s 24 ,000 31 ,219 48 ,722 65 ,050

Wing GroupIS , psf 8.4 8.9 9. 6 9.3 E q * GrplSelnpB psf 3 .8 4. 8 6.3 4. 2

U l t i m a t e Load F a c t o r , g * s 3.75* 3.75* 3.75* 3.75*


Wing, S 689 82 2 1 , 0 0 3 1 , 5 7 9 Horiz. T a i l , Sh 193 210 2 57 301

V e r t . T a i l , Sv 102 1 3 2 1 1 7 1 6 7

Empenn. Area , S =mp

295 342 3 74 46 8

*Assumed


Table A7.6a Group Weight Data f o r Turboprop. Transpor ts -- ~ , D P I I P Z I I P E I I P I I I I ~ P I P ~ ~ P P L I P I P I I . D ~ ~ ~ ~ P ~ I ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~

Br i s t o l Vickers TYPe - Br i t ann ia Canadair Viscount Lockheed

300 CL-44C 81 0 E l e c t r a Number of engines : 4 4 4 4 Weight Item, l b s


Nose Gear Main Gear


Engines A i r Induct. System F u e l System Prope l l e r I n s t . Propulsion System


Avionics + Instrum. Surface Controls ~ y d r a u l i c System Pneumatic System E l e c t r i c a l System Elec t ron ics APU - -

A i r Cond. System* 3,000 2,536 2,092 Anti- icing System - - Furnishings 6, 866 12,349 3,476 Auxil iary Gear o 0 0

Fixed Equipm't To ta l 15,082 22,788 10,505 14,469 ..................................

Max. Fuel Capacity 69,395 82,170 13,897 37,205 Payload (Max.) 30,000 37,630 15,054 18,907



T a b l e A7.6b Group Weight Data f o r Turboprop. T r a n s p o r t s

Br i s t o l V i c k e r s B r i t a n n i a C a n a d a i r V i s c o u n t

300 CL-4 4C 81 0

F l i g h t Des ign Gross Weight , GW, l b s 155 ,000 205,000 72 ,500

S t ruc tu re IGW 0.248 0.263 0.258 Power Plant/GW 0.128 0.111 F i x e d ~ q u i p m ' t / G ~ 0.097 0.111 0.145 Empty Weight IGW 0.587 0.516 0.569

Wing GroupIGW 0.087 0.077 0.086 Empenn. Group/GW 0.021 0 .018 0.017 F u s e l a g e GroupIG~ 0.072 0.100 0.095 Nacelle GroupIGW 0.032 0.033 0.025 Land. Gear GroupIGW 0.037 0.035 0.034

Take- off Gross Wht, WTO, l b s 155 ,000 205,000 72,500

Empty Weight , W E # Ibs 91,000 105 ,785 41,276

Lockheed E l e c t ra

Wing Group/S, psf 6.5 7.6 6.5 5.9 Emp. Grp/S emp* ps i 3.4 4.0 2.9 3.1

Ult imate Load F a c t o r , g ' s 3.75* 3.75* 3.75* 3.75.


Wing, S 2,075 2,075 963 1 , 3 0 0 Horiz. T a i l , Sh 588 588 307 399

V e r t . T a i l , S, 356 356 1 2 4 212


944 944 431 611

*Assumed


T a b l e A8. la Group Weight Data fo r Mi l i t a ry T r a i n e r s ~ ~ P ~ l a l P P I P P a % I ~ a P ~ 3 1 P I P P ~ L I P P P P ~ P a a P P ~ ~ a X = ~ 5 ~ = ~ ~ = = = ~

Rockwe l l Fouga Cana- Type N o r t h r o p NAA C e s s n a Magis- d a i r

T-38A T-39A T-37A t e r CL-41 Numbef of engines: 2 2 2 2 2 Weigh t Item, l b s

Wing Group Empennage Group F u s e l a g e Group E n g i n e S e c t i o n and. Gear Group

Nose Gear Main Gear


Eng i n e ( s A i r I n d u c t . Sys t em F u e l S y s t e m P r o p u l s i o n Sys t em


A v i o n i c s + Instrum. S u r f a c e C o n t r o l s H y d r a u l i c S y s t e m P n e u m a t i c Sys t em E l e c t r i c a l S y s t e m E l e c t r o n i c s A i r Cond. System** A n t i - i c i n g System F u r n i s h i n g s A u x i l i a r y Gear

765 1 , 7 5 3 5 3 1 1 , 0 8 9 892 305 297 1 2 8 1 6 5 2 0 1

1 , 9 8 5 2 ,014 83 9 743 955 147 315* i n fuse. 40 4 57 72 8 330 459 318

F i x e d Equ ipm ' t T o t a l 1 , 9 5 8 2 ,899 950 ..................................

Max. F u e l C a p a c i t y 3 , 9 1 6 5 ,805 1 , 9 5 9 1 , 2 9 9 2 ,082 P a y l o a d (Max. F u e l ) * * * 426 1 , 5 0 0 400 400 400

* N a c e l l e group f o r T-39A * * I n c l u d e s p r e s s u r i z a t i o n system * * * I n c l u d e s crew


T a b l e A8. lb Group Weight Data fo r Mili tary T r a i n e r s D P P I P S P I P P I P = = I P P I P = P = P P I L = P % P = = P P = P = P = = = = = = = = ~ = = = ~ ~

Rockwe l l Fouga Cana- TYPe N o r t h r o p NAA C e s s n a Magis- d a i r

T-3 8A T-39A T-37A te r CL-4 1

F l i g h t D e s i g n Gross Weight , GW, l b s 1 1 , 6 5 1 1 6 , 3 1 6 6 , 2 2 8 6 ,280 1 1 , 2 8 8

StructureIGW 0.314 0.313 0.294 0.391 0.213 Power Plant/GW 0.140 0.080 0 .192 F i x e d Equipmet/GW 0.168 0 .178 0 .152 N.A. N.A. Empty W t / G W 0.622 0.570 0 .638 0.755 0.576

Wing GrouplGW 0.066 0.107 0 .085 0.173 0.079 Emp. GroupIGW 0.026 0 .018 0 .021 0.026 0 .018 F u s e l . GroupIGW 0.170 0 .123 0.135 0 .118 0.085 E n g i n e Sect ionIGW 0.013 0.019* i n f u s . 0.004 Land. Gear Group/GW 0.039 0.045 0.053 0 .073 0 .028

Take- of f Gross W h t ~ WTOD l b s 1 1 , 6 5 1 1 6 , 7 0 1 6 ,436 6 ,280 1 1 , 2 8 8

Empty Weigh t , WE, l b s 7 ,247 9 , 3 0 7 3 ,973 4 ,740 5 ,296

Wing Group/S , psf 4.5 5 .1 3.9 5.9 4 .1 Emp. Grp/SempB psf 2.9 2.5 1 .7 3.4 3.4

Ultimate Load F a c t o r , 9 ' s 10.O** 6.0 10.0


Wing, S 170 342 1 3 5 186 220 Horiz. T a i l , Sh 59 77 54 * * * 41.3

V e r t . T a i l , S, 47.8 41.6 20.4 *** 17 .5


1 0 7 119 74.4 48. 8 58. 8

*Nacelle group f o r T-39A **Assumed

**V- t a i l


Table A9.la Group Weight Data f o r F igh te r s (USAF) P P l l l l l P P P I I I P a P x P P 1 P ~ I P = P . t P ~ = n I I P P P P I = a ~ ~ ~ ~ ~ ~ ~ = a

TYPe NAA McDonnell Gen. Dyn. F-100F F-lO1B RF-lO1C F-102A*

~umbeir of engines: 1 2 2 1 Weight Item, l b s

Wing Group 3,896 3,507 3,680 3,000 Empennage Group 979 81 2 83 7 535 Fuselage Group 4,032 3,901 3,955 3,409 Engine Sect ion 104 99 103 39 Land. Gear Group 1,509 1,592 1,596 1,056

Nose Gear Main Gear .................................

S t r u c t u r e T o t a l 10,520 9,911 10,171 8,039 ................................. Engine ( s ) 5,121 10,800 9,676 4,993 Air Induct. System 504 72 9 638 69 3 Fuel System 761 1,226 1,412 394 Propulsion System 414 892 599 27 8 ................................. Power P lan t Tota l 6,800 13,647 12,325 6,358 ................................. Avionics + Instrum. 303 318 204 141 Surface Controls 1,076 772 7 80 413 Hydraulic System 157 433 359 3 18 Pneumatic System E l e c t r i c a l System 568 82 5 81 9 594 Elec t ron ics 496 2,222 62 9 2,001 Armament 794 228 3 6 5 89 A i r Cond. System** 435 270 3 62 259 Anti- icing System Furnishings 427 4 80 242 2 27 Auxi l ia ry Gear 77 84 91 7 8 ................................. Fixed Equipm't To ta l 4,333 5,632 3,522 4,620 .................................

'oil+ 'ti0 166 223 223 216

Max. Fuel Capacity 7,729 8,892 9,782 7,053 Payload (Max. Fuel ) 250 1,881 704 1,241

*This a i r p l a n e is a d e l t a wing conf igura t ion **Includes p r e s s u r i z a t i o n system


T a b l e A9. lb Group Weight D a t a fo r F i g h t e r s (USAF) P I P I P 3 1 = I I P P P = I = P L E = P ~ = P n ~ I ~ = I t = = P = 1 = L ~ ~ I = n ~ = = n ~ ~

NAA ~ c D o n n e l 1 Gen.Dyn. F-100F F-1O1B RF-lO1C F-102Ae

F l i g h t D e s i g n Gross Weight , GW, l b s 29 ,391 39 ,800 3 7 , 0 0 0 2 5 , 5 0 0

S t r u c t u r e I G W 0 .358 0 ,249 0.275 0.315 Power Plant/GW 0.231 0.343 0.333 0 .249 F i x e d Equiprn't/GW 0.147 0.142 0.095 0 .181 Empty Weight/GW 0.737 0.733 0.724 0.750

Wing GrouplGW 0.133 0 .088 0.099 0 .118 Empenn. Group/GW 0.033 0.020 0.023 0 .021 F u s e l a g e GroupIGW 0 ,137 0 .098 0.107 0.134 E n g i n e Sect ionIGW 0.004 0.002 0 .003 0.002 Land. G e a r GrouplGW 0 .051 0 , 0 4 0 0.043 0.041

Take- of f Gross Wht, WTO, l b s 30 ,638 4 1 , 2 8 8 37 ,723 28 ,137

Empty Weigh t ,

W ~ ' l b s 21 ,653 29 ,190 26,774 1 9 , 1 3 0

Wing GroupIS , p s f 9.7 9.5 10.0 4.3 Emp. G r p l S emp, psf 6.3 5.1 5 ,2 5.6

U l t i m a t e Load F a c t o r , g ' s 7.33 10.2 11.0 10.5


Wing, S 400 368 3 68 69 8 Horiz. T a i l , Sh 98. 9 75.1 75. 1 0

V e r t . T a i l . S, 55.6 84.9 84.9 95.1

Empenn. Area , S 1 5 5 1 6 0 1 6 0 95.1 emp

T h i s airplane is a d e l t a wing conf igura t ion

Part V Appendix A P a g e 1 6 5

Table A9.2a Group Weight Data f o r F i g h t e r s (USAF) PXPPl=PPIPPaPIP=IP=~=======PI=5==a=PPP=P~=a===Pma

TYPe Republic Gen.Dyn. North American F-lO5B F-106A* F-107A F-86H

~umbe? of engines : 1 1 1 1 Weight I t e m , l b s

Wing Group 3 , 4 0 9 3 , 3 0 2 3 , 7 8 7 2 , 7 0 2 Empennage Group 9 6 5 6 9 3 1 , 1 3 0 3 2 9 Fuselage Group 5 , 870 4 , 4 0 1 4 , 7 9 2 2 , 0 3 5 Engine S e c t i o n 1 0 6 39 2 6 0 4 2 Land. Gear Group 1 , 8 4 8 1 , 2 3 2 1 , 4 1 0 9 89


S t r u c t u r e T o t a l 1 2 , 1 9 8 9 , 6 6 7 1 1 , 3 7 9 6 , 0 9 7 ................................. Engine 6 , 1 8 7 5 , 8 1 6 6 , 1 0 0 3 , 6 4 6 A i r Induct . System 5 2 4 975 83 3 1 6 7 Fuel System 6 0 8 7 7 7 9 83 84 5 Propuls ion System 4 0 6 5 0 3 3 68 3 4 0 ................................. Power P l a n t T o t a l 7 , 7 2 5 8 , 0 7 1 8 , 2 84 4 , 9 9 8 ................................. Avionics + Instrum. 2 27 1 9 0 2 8 8 111 Sur face Con t ro l s 1 , 3 1 1 4 4 5 1 , 4 5 4 3 5 8 Hydraulic System Pneumatic Svstem E l e c t r i c a l gystem 7 0 0 6 0 6 4 47 47 6 E l e c t r o n i c s 7 3 7 2 , 7 4 3 3 82 2 3 0 Armament 719 626 1 , 0 0 6 82 8 Air Cond. Systema* Ant i- ic ing System 1 6 8

- - Furn ish ings 2 4 3 2 9 0 2 82 1 8 2 Auxi l i a ry Gear 9 2 69 4 2 1 2 APU 2 2 4 0 0 0 ................................. Fixed Equipm't T o t a l 4 , 8 7 0 5 , 8 0 7 4 , 4 4 1 2 , 7 4 1 .................................

Max. Fuel Capaci ty 7 , 5 4 0 8 , 4 7 6 1 1 , 0 5 0 3 , 6 6 0 Payload (Max. F u e l ) 7 5 7 1 , 3 7 4 2 , 5 6 0 4 2 0

*This a i r p l a n e is a d e l t a wing c o n f i g u r a t i o n **Includes p r e s s u r i z a t i o n system


T a b l e A9.2b Group Weight Data f o r F i g h t e r s (USAF) ~ P = ~ = = = D I I P P P P P I P I P ~ ~ P L : P ~ P P X P P L : P P E P P ~ P = ~ ~ P ~ ~ ~ ~ ~ ~ ~

R e p u b l i c Gen.Dyn. N o r t h American F-IOSB F-106A* F-107A F-86H

F l i g h t Design Gross Weight , GW, l b s 31 ,392 30 ,590 29 ,524 19 ,012

S t r u c t u r e I G W 0.3 89 0.316 0 .385 0 .321 Power Plant lGW 0.246 0.264 0 , 2 8 1 0.263 F i x e d E q u i p m ' t I G ~ 0.155 0 ,190 0.150 0.144 Empty Weight IGW 0.797 0.766 0. 816 0 .728

Wing GroupIGW 0.109 0.108 0 .128 0 .142 Empenn. GrouplGW 0 .031 0 .023 0 .038 0.017 F u s e l a g e GroupIGW 0.187 0.144 0.162 0.107 E n g i n e SectionIGW 0.003 0 .001 0.009 0.002 Land. Gear GroupIGW 0.059 0.040 0 .048 0.052

Take- of f Gross Wht, WTO, l b s 34 ,081 3 3 , 8 8 8 39 ,405 1 8 , 9 0 8

Empty Weight,

W ~ ' l b s 25 ,022 2 3 , 4 4 8 24 ,104 1 3 , 8 3 6

Wing GroupIS , psf 8.9 4.7 9.6 8.6 Emp. G r p I S emp, p ~ f 5.2 6.6 6.4 4 .1

U l t i m a t e Load F a c t o r , g ' s 13 .0 10.5 13 .0 11 .0


Wing, S 3 85 69 8 39 5 313 Horiz. T a i l , Sh 96.5 o 93.3 47.2

V e r t . T a i l , S, 88. 1 1 0 5 83. 8 32.2


1 8 5

T h i s airplane is a d e l t a wing conf igura t ion


Table A9.3a Group Weight Data f o r F igh te r s (USN) ~ P D ~ ~ o I I ~ ~ ~ ~ ~ ~ ~ P I I I P ~ P O P ~ L P I O ~ % % I I I P ~ ~ P P ~ ~ ~ ~ ~ ~ ~ %

TYPe Vought McDonnell Grumman F 813-3 F4H FllF F9F-5

Number* of engines : 1 2 1 1 Weight I t e m , l b s

Wing Group 4,128 4,343 2,180 2,294 Empennage Group 1,045 853 6 69 404 Fuselage Group 3,850 4,042 3,269 1,779 Engine Sec t ion 92 125 47 o Land. Gear Group 949 1,735 907 72 8


S t r u c t u r e To ta l 10,064 11,098 7,072 5,205 ................................. E n g i n e ( ~ ) 6,010 6,940 3,4 89 2,008 A i r Induct. System 673 1,037 159 225 Fuel System 84 9 953 463 529 Propulsion System. 3 3 8 106 192 116 ................................. Power P l a n t To ta l 7,870 9,036 4,303 2,878 ................................. Avionics + Instrum. 191 166 118 82 Surface Controls 1,425 919 760 345 Hydraulic System Pneumatic System E l e c t r i c a l system 439 502 459 458 Elec t ron ics 84 0 1,386 439 292 Armament 376 446 35 8 416 A i r Cond. System* Anti- icing System - - Furnishings 210 321 166 144 Auxi l ia ry Gear 183 0 131 51

Fixed Equipm't To ta l 4,143 4,522 2,673 2,140

Max. Fuel Capacity 14,306 13,410 6,663 7,160 Payload (Max. Fuel ) 1,197 1,500 340



T a b l e A9.3b Group Weight D a t a fo r F i g h t e r s (USN) P I I P I I P P I P P I E I I P E P P ~ D I P P P E I : ~ ~ ~ ~ P . ~ E ~ P ~ . C P P ~ E ~ ~ P U ~ ~

Vought McDonnell Grumman F8U-3 F4H F l l F F9F-5

F l i g h t Design Gross Weigh t , GW, l b s 3 0 , 5 7 8 3 4 , 8 5 1 1 7 , 5 0 0 1 4 , 9 0 0

S t r u c t u r e / G W 0.329 0.318 0.404 0.349 Power Plant/GW 0.257 0.259 0 .246 0.193 F i x e d EquipmPt/GW 0.135 0.130 0.153 0.144 Empty Weight/GW 0.722 0.707 0.771 0.686

Wing Group/GW 0.135 0.125 0.125 0.154 Empenn. Group/GW 0.034 0.024 0.03 8 0.027 F u s e l a g e Group/GW 0.126 0 .116 0.187 0.119 E n g i n e Sect ion/GW 0.003 0.004 0.003 0 Land. Gear Group/GW 0.031 0.050 0.052 0.049

Take- of f Gross Wht, WTO, l b s 3 8 , 5 2 8 40 ,217 21 ,233 1 7 , 5 0 0

Empty Weight, W E B l b s 22 ,092 24 ,656 1 3 , 4 8 5 1 0 , 2 2 3

Wing Group/S , psf 8.9 8.2 8. 5 9.2 Emp. Grp/Semp, psf 7.2 5.2 5. 8 3.5

U l t i m a t e Load F a c t o r , g ' s 9.6


Wing, S 462 530 255 250 Horiz. T a i l , Sh 67.2 96.2 65.5 4 8

V e r t . T a i l , S, 78.6 67.5 50.3 6 6

Empenn. Area , Semp 1 4 6 164 1 1 6 114


Table A 9 . 4 a Group Weight Data. f o r F i g h t e r s (USN) ==P==PP915==PXXXPPP=I=============L:==~=P==a=====

TYPe Grumman McDonnell NAA Vought 3 A2F(A6) F3H-2 A 3 J F7U-1

Number of eng ines : 2 2 2 1 Weight Item, l b s

Wing Group 4 , 7 3 3 4 , 3 1 4 5 , 0 7 2 3 , 5 8 3 Empennage Group 81 9 5 7 6 1 , 3 5 8 7 2 6 Fuselage Group 3 , 5 3 8 3 , 5 5 1 6 , 8 5 1 9 3 7 Engine S e c t i on 64 9 3 80 Land. Gear Group 2 , 3 4 3 1 , 4 5 8 2 , 1 7 3 1 , 1 8 1


S t r u c t u r e T o t a l 1 1 , 4 9 7 9 , 9 9 2 1 5 , 5 3 4 6 , 4 2 7

Engine(s ) 4 , 0 1 0 4 , 9 6 0 7 , 2 6 0 2 , 7 9 0 Air Induct . System 6 1 6 1 4 7 6 7 69 0 Fuel System 9 3 6 1 , 2 6 2 9 7 9 1 , 0 8 0 Propuls ion System. 6 3 2 7 0 3 5 3 9 3 7 ................................. Power P l a n t T o t a l 5 , 6 3 9 6 , 9 0 6 9 , 3 5 9 5 , 4 9 7 ................................. Avionics + Instrum. 1 3 3 1 4 5 2 1 0 1 0 8 Sur face Con t ro l s 9 3 2 1 , 0 6 7 1 , 8 4 5 4 82 Hydraulic System Pneumatic Svstem E l e c t r i c a l gystem 69 5 5 3 5 82 1 3 7 1 E l e c t r o n i c s 2 , 6 5 2 9 8 4 2 , 2 3 9 3 2 8 Armament 3 2 3 6 6 2 4 5 3 6 7 A i r Cond. System* Ant i- ic ing System

- -

Furn ish ings 4 7 6 2 1 8 6 7 6 2 7 9 Auxi l i a ry Gear 2 5 3 1 2 8 ................................. Fixed Equipm't T o t a l 5 , 5 4 5 4 , 4 3 9 6 , 5 3 5 2 , 4 5 9 .................................

Max. Fue l Capaci ty 8 , 7 6 4 9 , 7 8 9 1 9 , 0 7 4 5 , 8 2 6 Payload (Max. F u e l ) 2 , 0 0 0 2 1 6 1 , 8 8 5 5 0 2



T a b l e A9.4b Group Weight Data f o r F i g h t e r s (USN) 5PPPltPPlllltPI%=PP~PPIPP=P~PPalPlP=PPP~=xl=~Pu~

Grumman McDonnel 1 NAA Vought A2F(A6) F3H-2 A3J F7U-I*

F l i g h t D e s i g n Gross Weight . GW, l b s 34 ,815 26 ,000 4 6 , 0 2 8 1 9 , 3 1 0

S t r u c t u r e I G W 0.330 0.3 84 0.337 0.333 Power Plant/GW 0.162 0 .266 0 .203 0.285 F i x e d Equipm0t/GW 0.159 0.171 0.142 0.127 Empty Weight/GW 0 .651 0. 818 0.679 0.746

Wing Group/GW 0.136 0.166 0.110 0.186 Empenn. Group/GW 0.024 0.022 0.030 0.03 8 F u s e l a g e Group/GW 0.102 0.137 0 ,149 0 .048 E n g i n e Section/GW 0.002 0.004 0.002 Land. Gear Group/GW 0.067 0.056 0.047 0 .061

Take- of f Gross W h t ~ WTOD l b s 34 ,815 32 ,037 53 ,658 21 ,638

Empty Weigh t , WE, I b s 22,680 21 ,272 31,246 14 ,397

Wing GroupIS , p s f 9.1 8.4 7.2 7.1 Emp. GrpISempD p s i 4.4 4.5 3.4 8.3

U l t i m a t e Load F a c t o r , g ' s


Wing, S 520 516 700 507 Horiz. T a i l , Sh 1 2 0 82.5 304 O*

V e r t . T a i l , Sv 68.4 45.4 1 0 1 88


1 8 8 1 2 8 405 88

* T h i s airplane is es sen t i a l ly a f l y i n g wing w i t h two ver t ica l t a i l s


Table A9.5a Group Weight Data f o r F igh te r s (USAF and U S N )

Type McDonnell Douglas - F-4E F-1 SC F/A-18A AV-8B* ( USAF 1 ( USAF ) ( U S N ) ( U S N 1


Wing Group 5,226 3,642 3,798 1,443 Empennage Group 969 1,104 945 372 Fuselage Group 5,050 6,245 4,685 2,060 Engine Sect ion 166 102 143 141 Land. Gear Group 1,944 1,393 1,992 1,011

Nose Gear 377 2 64 62 6 334 Main Gear 1,567 1,129 1,366 400 Outrigger Gear 277 ...................................

S t r u c t u r e To ta l 13,355 12,486 11,563 5,027 ................................... Engine ( s ) 7,697 6,091 4,294 3,815 A i r Induct. System 1,318 1,464 423 236 F u e l System 1,932 1,128 1,002 542 Propulsion System. 312 522 558 444 ................................... Power P l a n t To ta l 11,259 9,205 6,277 5,037 ................................... Instrument group 270 151 94 80 Surface Controls 1,167 81 0 1,067 69 8 Hydraulic System 543 433 3 64 176 Pneumatic System E l e c t r i c a l System 542 607 5 47 424 Elec t ron ics 2,227 1,787 1,538 697 Armament 64 1 627 3 87 152 Air Cond. System 406 685 5 93 218 P r e s s u r i z a t i o n Syst. Anti- icing System 21 Furnishings 611 294 317 298 Auxi l ia ry Gear 412 119 189 Photographic System 24 B a l l a s t 318 3 6 Manuf. Var ia t ion 57 -97 -19 -1 6 ................................... Fixed Equipm't T o t a l 6,900 5,734 5,134 2,727 ................................... Max. Fuel Capacity 12,058 13,455 17,592** 7,759 Expendable Payload 2,193 2,571 5,453 4,271 Fixed Payload*** i n c l . i n armament 2,231 83 2

*V/STOL f i g h t e r **Incl. 6,732 l b s ex t . f u e l ***Pylons, racks, launchers , FLIR and camera pods


T a b l e A9.5b Group Weight Data for F i g h t e r s (USAF a n d USN) I I P D P ~ P I D P ~ ~ P P ~ ~ P P D % P R ~ C P P P P E P P ~ P ~ I C % P I E P . C E P P ~ ~ ~ ~ ~ E I P ~ ~ E ~ ~ ~

McDonnell D o u g l a s F-4E F-15C F/A-18A AV-8B*

F l i g h t D e s i g n Gross Weight, GW, l b s 37 ,500 37 ,400 32 ,357 22 ,950

S t r u c t u r e / G W 0.356 0.334 0.357 0.219 Power Plant/GW 0.300 0.246 0.194 0.219 F i x e d Equiprn't/GW 0.182 0.147 0 .158 0.120 Empty Weight /GW 0.840 0.733 0.710 0.557

wing Group/GW 0.139 0.097 0.117 0.063 Empenn. Group/GW 0.026 0.030 0.029 0.016 F u s e l a g e Group/GW 0.135 0.167 0.145 0.090 E n g i n e Sect ion/GW 0.004 0.003 0.004 0.006 Land. Gear GrouplGW 0.052 0.037 0.062 0 .044

Take- of f Gross Wht. WTO, I b s 5 8 ,000 68 ,000 51 ,900 29 ,750

Empty Weight, W E D l b s 31 ,514 27 ,425 22,974 1 2 , 7 9 1

Wing Group/S , psf 9.5 6.1 9.5 6.3 Emp. Grp /S emps psf 5 .8 4.7 4.9 5.0

Ultimate Load F a c t o r , g ' s 9.75 11.0 11.25 10.5


Wing, S 5 4 8 599 400 230 Horiz. T a i l . Sh 100 111 88.1 48. 5

V e r t . T a i l , S, 67.5 1 2 5 104 26.6


1 6 8 236 1 9 2 75.1

*V/STOL F i g h t e r

P a r t v *-

Appendix A P a g e 1 7 3

Table A l O . l a Group Weight Data f o r M i l i t a r y Je t = = ~ L ~ t = = = P l E = P P E P P P P m a = 5 . C = 5 5 = L 5 5 I C O = = = = I S = = = = = = = =

Transpor t s =5====st3=:

* TYPe Boeing Lockheed

KC135* C- 141B C-5A Number of engines: 4 4 4 Weight Item, l b s

Wing Group 2 5 , 2 5 1 3 5 , 2 7 2 1 0 0 , 0 1 5 Empennage Group 5 , 0 7 4 5 , 9 0 7 1 2 , 4 6 1 Fuselage Group 1 8 , 8 6 7 3 6 , 8 5 7 1 1 8 , 1 9 3 Nacel le Group 2 , 5 7 5 5 , 1 6 8 9 , 5 2 8 Land. Gear Group 1 0 , 1 8 0 1 0 , 8 5 0 3 8 , 3 5 3

Nose Gear 1 , 2 3 4 4 , 4 5 5 Main Gear 9 , 6 1 6 3 3 , 8 9 8 .........................

S t r u c t u r e T o t a l 6 1 , 9 4 7 9 4 , 0 5 4 2 7 8 , 5 5 0 ......................... E n g i n e ( ~ ) 1 6 , 6 8 7 2 3 , 6 6 5 3 8 , 0 3 5 A i r Induct . System 1 7 2 Fuel System 4 , 0 5 2 1 , 8 0 2 2 , 5 4 0 Propuls ion System 5 9 1 ......................... Power P l a n t T o t a l 2 1 , 5 0 2 2 5 , 4 6 7 4 0 , 5 7 5 ......................... Avionics + Instrum. 553 3 , 0 7 8 3 , 8 2 3 Sur face Con t ro l s 2 , 0 4 4 3 , 7 0 1 7 , 4 0 4 Hydraulic System 85 8 1 , 6 0 4 4 ,086 Pneumatic System Electrical System 2 , 4 7 0 2 , 8 2 6 3 , 5 0 3 E l e c t r o n i c s 2 , 0 9 6 1 , 1 6 3 9 9 2 APU 0 534 9 87 Oxygen System 4 7 9 3 0 8 Air Cond. Systeme* 1 , 4 6 4 2 , 2 8 3 3 , 4 1 6 Ant i- ic ing System 4 5 3 2 2 9 Furn i sh ings 1 , 5 1 8 5 , 2 1 0 1 9 , 2 7 2 A u x i l i a r y Gear 1 , 8 9 9 1 0 3 3 9 ......................... Fixed Equipm't T o t a l 1 2 , 9 0 2 2 1 , 4 3 4 4 4 , 0 5 9

Max. Fue l Capaci ty 1 5 8 , 9 9 7 1 5 3 , 3 5 2 3 3 2 , 5 0 0 Payload (Max.) 7 3 , 8 7 3 2 0 0 , 0 0 0

*This is a t a n k e r a i r p l a n e **Includes p r e s s u r i z a t i o n system

p a r t v Appendix A Page 174

T a b l e AlO. lb Group Weight Data fo r M i l i t a r y Je t

T r a n s p o r t s ======rn===

Boeing KC135*

F l i g h t D e s i g n Gross Weigh t , GW, l b s 297 ,000

S t r u c t u r e I G W 0.209 Power PlantIGW 0.072 F i x e d Equipm'tIGW 0.043 Empty Weight/GW 0.323

Wing GroupIGW 0.085 Empenn. GroupIGW 0.017 F u s e l a g e GroupIGW 0.064 Nacelle GrouplGW 0.009 Land. Gear GroupIGW 0.034

Take- of f Gross Wht , WTOn I b s 297 ,000

Empty Weight, WE, I b s 95 ,938

Lockheed C-141B C-5A

Wing GroupIS, psf 10.4 10.9 1 6 . 1 Emp. GrplSemp, psf 6.2 6.6 6.5

Ultimate Load F a c t o r , g ' s 3.75 3.75 3.75**


Wing, S 2 ,435 3 , 2 2 8 6 ,200 Horiz. T a i l . Sh 500 4 83 966

V e r t . T a i l , S, 312 416 961


81 2 89 9 1 , 9 2 7

* T h i s is a t a n k e r airplane * * a f t e r 100 ,000 l b s of f u e l h a s b e e n u s e d .


Table A10.2a Group Weight Data f o r TurbofPropel ler S = = P ~ P I P P P P P I I P ~ P I I P P ~ P P ~ P P I : ~ : ~ P P P P P P I ~ ~ ~ P ~ ~ ~ I ~ ~ ~ ~ ~ ~

Driven Mi l i t a ry Transports =PPXXPPPPLPPI I I==PP~PPILe=P

A.W.(HS) Douglas Lockheed Breguet Argosy C-133A C-130H 941*


Wing Group 10,800 27,403 13,950 4,096 Empennage Group 1,300 6,011 3,480 1,387 Fuselage Group 11,100** 30,940 14,695 6,481 Nacelle Group 1,200 3,512 2,756 i n wing Land. Gear Group 3,180 10,635 5,309 2,626

Nose Gear 730 Main Gear 4,579 ..................................

S t r u c t u r e To ta l 27,580 78,501 40,190 14,590 .................................. Engines 10,470 13,746 A i r Induct. System Fuel System 1,338 3,105 Prope l l e r I n s t . 5,403 i n eng. Propulsion System 2,081 i n eng. .................................. Power P lan t To ta l 19,292 16,851 .................................. Avionics + Instrum. 57 8 3,582 Surface Controls i n s t r u c t . 1,804 1,673 1,056 Hydraulic System 2,678 6 64 Pneumatic System E l e c t r i c a l System 2,004 2,459 Elec t ron ics 2,047 i n av ionics APU 188 651 Oxygen System 231 A i r Cond. System*** 2,973 1,684 Anti- icing System 797 .

Furnishings 3,632 4,472 Auxil iary Gear 117 6 Operating items 532 .................................. Fixed Equipm't To ta l 16,021 16,219 ..................................

Max. Fuel Capacity Payload (Max.

*This is a STOL a i r p l a n e **Tailbooms a t 2,360 l b s a r e included ***Includes p r e s s u r i z a t i o n system


T a b l e A10.2b Group Weight D a t a f o r T u r b o l P r o p e l l e r PaPOwPPIIIIllt=t~EPIIPtIIL=IIPIII=IIIPIIII~IIIIn~~

Driven Mi l i t a ry T r a n s p o r t s nPIIIIPII~PIIII~IIIIIm=P=u

A.W.(HS) D o u g l a s Lockheed B r e g u e t Argosy C-133A C-130H 9 4 1

F l i g h t D e s i g n Gross Weigh t , GW, l b s 82,000 275 ,000 155 ,000 5 8 , 4 2 1

S t r u c t u r e I G W 0.336 0.2 85 0.259 0 .250 Power P l a n t /GW 0.070 0.109 F i x e d Equipm' t /GW 0 .058 0.105 Empty WeightIGW 0.561 0.414 0.473 0 .508

Wing GroupIGW 0.132 0.100 0.090 0.070 Empenn. GroupIGW 0.016 0.022 0.022 0.024 F u s e l a g e GroupIGW 0.135 0.113 0.095 0:111 Nacelle Group/GW 0.015 0.013 0 .01 8 i n wing Land. G e a r GroupIGW 0.039 0.039 0.034 0 .045

Take- of f G r o s s Wht, WTO, l b s 82 ,000 275 ,000 1 5 5 , 0 0 0 5 8 , 1 2 1

Empty Weigh t , W E D l b s 46 ,000 113 ,814 73 ,260 29 ,675

Wing GroupIS , psf 7.4 10.3 8.0 4.5 Emp. G r p I S

emp' psf 2.3 4.2 4.2 2.6

U l t ima te Load F a c t o r , g ' s 3.75* 2 .50 3-75. 3-75.


Wing, S 1 , 4 5 8 2 ,673 1 , 7 4 5 902 Horiz. T a i l . Sh 327 80 1 536 320

V e r t . T a i l . S, 250 64 1 300 223


577 1 , 4 4 2 83 6 5 43


Table A10.3a Group Weight Data f o r P i s t o n / P r o p e l l e r = = t l P P P I I I = P P 1 P I ~ P I ~ = P f = P P P P . 5 = P = P P P ~ ~ = S 5 ~ = ~ S = P P ~ P = ~

Driven Mi l i t a ry Transpor ts POtPPPPtPPeOPPPPPPPPPU===P

* TYPe Beech Chase DeHavill. F a i r c h i l d

L-23F* C-123B Caribou C-119B Number of engines: 2 2 2 2 Weight Item, l b s

Wing Group 692 6,153 2,925 7,226 Empennage Group 156 1,103 790 1,193 Fuselage Group 679 7,763 2,849 7,157 Nacelle Group 279 1,633 7 81 2,538.e Land. Gear Group 4 53 2,081 1,230 4,197

Nose Gear Main Gear ................................

S t r u c t u r e To ta l 2,259 18,733 8,575 22,311 ................................ Engines 1,015 4,810 3,170 6,500 Prope l l e r s 260 1,430 87 1 Fuel System 127 671 221 Propulsion System 215 1,103 4 53 ................................ Power P l a n t To ta l

Avionics + Instrum. Surface Controls Hydraulic System Pneumatic System E l e c t r i c a l System Elec t ron ics APU A i r Cond. System Anti- icing System Furnishings Auxil iary Gear

Fixed Equipm't Tota l

Water 0 225 0 Max. Fuel Capacity 1,080 5,452 15,540 Payload (Max. 1,090 16,000 7,344

*Mil i ta ry ve r s ion of Twin Bonanza **Tailbooms included


T a b l e A10.3b Group Weight Data for P i s t o n / P r o p e l l e r O = D P P O I P P P P P P I I l P I I P P P P P P P = P P I P = ~ . C P P P = P P = s 6 P = = = ~ = E =

Driven Mil i ta ry T r a n s p o r t s P I P P E I P O P P P = I P P P I ~ ~ P = = P ~ P P

Beech C h a s e DeHavill. F a i r c h i l d L-23F* C-123B C a r i b o u C-119B

F l i g h t Design Gross Weigh t , GW, l b s 7 , 3 6 8 54 ,000 26 ,000 64 ,000

S t r u c t u r e / G W 0.307 0.347 0.330 0.349 Power Plant/GW 0.219 0 .148 0 .181 0.187 F i x e d Equipm't/GW 0.144 0.062 0.062 0.107 Empty Weight /GW* 0.670 0 .558 0 .635 0.641

Wing Group/GW 0.094 0.114 0 .113 0 .113 Empenn. Group/GW 0.021 0.020 0 .030 0.019 F u s e l a g e Group/GW 0.092 0 .144 0 .110 0.112 Nacelle GrouplGW 0.038 0.030 0.030 0.040** Land. Gear Group/GW 0 .061 0.039 0.047 0.066

Take- of f Gross W h t ~ WTO, I b s 7 , 3 6 8 52 ,802 26 ,000 64 ,000

Empty Weight, WE, I b s 4 ,936 30 ,108 1 6 , 5 0 0 41 ,017

Wing Group/S , psf 2.5 5.0 3.2 5.0 Emp. Grp/Semp. psf 1.7 2 .1 1 .9 2.7

U l t i m a t e Load F a c t o r , g ' s 6.6 4.5 5.1


Wing, S 277 1 , 2 2 3 912 1 ,447 Horiz. T a i l , Sh 29.3 206 2 97

V e r t . T a i l , Sv 65.1 2 1 1 1 5 1


94.4 520 417 44 8

* T h i s is a mi l i t a ry ve r s ion of t h e Twin Bonanza **Tai lbooms i n c l u d e d


Table A10.4a Group Weight Data f o r P i s ton tPrope l l e r D p l f l P 3 P P P P 3 P I I P I P P P I C P O P = P P I I I P P P P = P P P 3 5 = = U = = 5 = = = = = =

Driven M i l i t a r y Transports =r==a==P=P==PPP=I==P===P=P

i

TYPe Douglas Boeing Lockheed C-124c C-97c C- 69 C-121A


Wing Group 1 8 , 1 3 5 1 5 , 3 89 9 , 4 6 6 1 1 , 1 8 4 Empennage Group 3 , 0 2 5 2 , 0 7 8 2 , 0 2 6 2 , 0 9 4 Fuselage Group 1 8 , 0 7 3 1 3 , 5 7 2 6 , 7 9 4 8 ,520 Nacel le Group 6 , 1 1 9 4 , 7 5 3 2 , 5 0 5 3 , 9 7 0 Land. Gear Group 1 1 , 7 0 1 7 , 1 1 2 4 , 4 8 1 4 , 7 7 1

Nose Gear 963 1 , 0 1 9 1 , 0 7 7 Main Gear 6 , 1 4 9 3 , 4 6 2 3 , 6 9 4 ..................................

S t r u c t u r e Tota l 57,053 4 2 , 9 0 4 2 5 , 2 7 2 30 ,539 .................................. E n g i n e ( ~ ) 1 5 , 5 5 1 1 3 , 8 4 4 1 0 , 5 6 8 1 1 , 5 3 6 Air Induct. System 4 , 0 4 6 Fuel System 4 , 0 5 9 Prope l l e r Ins t . 4 , 3 6 3 Propulsion System i n prop. .................................. Power P lan t T o t a l 2 8 , 0 1 9 2 3 , 0 5 1 1 5 , 6 3 3 1 5 , 6 7 6 .................................. Avionics + Instrum. 7 6 9 Surface Controls 1 , 4 9 3 Hydraulic System Pneumatic System 5 82

E l e c t r i c a l System 1 , 9 5 2 Elec t ron ics 1 , 8 8 6 APU 4 1 0 Air Cond. System* Anti- icing System 3 , 2 9 4

Furnishings 7 , 5 3 9 Auxil iary Gear 1 0 4 .................................. Fixed Equipm't T o t a l 1 8 , 0 2 9 9 , 9 9 7 7 , 6 2 5 1 3 , 7 1 0 .................................. Wail+ W t f o 3 , 3 8 9

Water and Alcohol 522 Max. Fuel Capacity 2 2 , 0 0 0 2 3 , 0 9 4 3 3 , 4 7 0 4 1 , 4 9 6 Payload (Max.) 5 5 , 2 6 2 4 6 , 5 0 0 1 2 , 3 3 0 1 2 , 5 5 0



T a b l e A10.4b Group Weight Data f o r P i s t o n l P r o p e l l e r I P E P I P I P P P ~ P P ~ I I P ~ ~ ~ . I I ~ P P ~ ~ I ~ P P ~ P P ~ I L ~ P R R I I ~ ~ ~ ~ ~ ~ ~ ~ ~ ~

Driven Mil i ta ry T r a n s p o r t s ~IPPI~PIPPPPEIIP~I~P~PPPPP

Douglas Boe ing Lockheed C-124C C-97C C- 69 C-121A

F l i g h t Design Gross Weigh t , GW, l b s 185 ,000 1 5 0 , 0 0 0 82 ,000 132 ,800

S t r u c t u r e / G W 0.308 0.286 0 .308 0.230 Power PlantIGW 0 .151 0.154 0.191 0 .118 F i x e d Equipmtt/GW 0.097 0.067 0.093 0.103 Empty Weight/GW 0.552 0.506 0.592 0.450

Wing Group/GW 0 .098 0.103 0.115 0.084 Ernpenn. Group/GW 0.016 0.014 0.025 0.016 F u s e l a g e Group/GW 0 .098 0.090 0.083 0.064 Nacelle Group/GW 0.033 0.032 0 .031 0.030 Land. Gear Group/GW 0.063 0.047 0.055 0.036

Take- off G r o s s Wht, WTO, I b s 185 ,000 1 5 0 , 0 0 0 82,000 132 ,800

Empty Weight, WE, l b s 102 ,181 75,974 48 ,530 59 ,715

Wing Group lS , psf 7.2 8.7 5.7 6. 8 Emp. G r p / S

emp' psf 2.6 3.3 2.9 3.0

u l t i m a t e Load F a c t o r , g 8 s 3.75 3.75 3.75 3.75


Wing, S 2 , 5 0 6 1 , 7 6 9 1 , 6 5 0 1 , 6 5 0 Horiz. T a i l , Sh 6 81 333 464 464

V e r t . T a i l , Sv 465 306 262 242

Empenn. Area , S ew 1 , 1 4 6 639 706 70 6


Table A1O.Sa Group Weight Data f o r M i l i t a r y P a t r o l ~ = ~ = I P ~ 0 3 1 S P 1 1 3 1 P P O P ~ = ~ P P = ~ ~ 5 i = ~ = O ~ ~ ~ P P = : = ~ ~ 5 = = = = ~ ~ =

A i rp l anes ==as======

'I

TYPe Grumman Lockheed Lockheed S2F- 1 P2V-4 U2

Number of engines: 2 2 1 Weight Item, l b s P i s t o n / P r o p e l l e r J e t

Wing Group 2 , 9 0 2 7 , 4 9 8 2 , 0 3 4 Empennage Group 6 81 1 , 5 8 9 3 2 0 Fuselage Group 1 , 7 0 1 5 , 1 5 5 1 , 4 1 0 Nacelle Group 9 6 5 2 , 3 0 3 0 Land. Gear Group 1 , 3 9 6 3 , 7 1 5 2 6 3

Nose Gear t a i l g e a r 6 0 Main Gear 2 0 3 .........................


Engines P r o p e l l e r s Fuel System Propuls ion System


Avionics + Instrum. Su r face Con t ro l s Hydraulic System Pneumatic System E l e c t r i c a l System 9 8 8 1 , 5 0 3 2 9 0 E l e c t r o n i c s 2 , 3 1 0 2 , 9 0 3 1 6 6 Armament 2 5 6 1 , 7 0 5 A i r Cond. System* Ant i- ic ing System - - Furn ish ings 65 7 1 , 3 2 7 8 2 Auxi l i a ry Gear 2 81 0 1 9 3

Fixed Equipm't T o t a l 5 , 9 1 7 9 , 3 7 2 1 , 3 5 1

Engine o i l 1 2 0 Armament P rov i s ions 1 8 5 , 9 5 1 Water 0 1 , 4 8 0 Max. Fuel Capaci ty 3 , 1 2 6 1 4 , 0 0 6 5 , 8 1 0 Payload (Max. 1 , 9 3 8 5 1 8 Crew N. A. N. A. 2 85

*Inc l . p r e s s . system **Inc l . a i r i nduc t ion and exhaus t s


T a b l e A10.5b Group Weight Data f o r Mi l i t a ry P a t r o l S ~ S P = I ~ S S P P P P P P P P P P P O P . L P ~ ~ P ~ P P P P ~ S P P O P P O U ~ U P ~ ~ ~ E ~ . L

A i r p l a n e s = = P . i s P P I P

TYPe Grumman Lockheed Lockheed S2F-1 P2V-4 U2 P i s t o n / P r o p e l l e r Je t

F l i g h t Design Gross Weigh t , GW, l b s 23 ,180 67 ,500 1 7 , 0 0 0

S t r u c t u r e / G W 0.330 0.300 0.237 Power Plant/GW 0 .191 0 .168 0.286 F i x e d Equipmet/GW 0.255 0.139 0.079 Empty Weight/GW 0.775 0.607 0 .603

Wing Group/GW 0.125 0 .111 0.120 Empenn. Group/GW 0.029 0.024 0.019 F u s e l a g e GrouplGW 0.073 0.076 0.083 Nacelle Group/GW 0.042 0.034 Land. Gear Group/GW 0.060 0.055 0.015

Take- off Gross Wht, WTO, l b s 24 ,167 67 ,500 1 9 , 9 1 3

Empty Weigh t , WE, I b s 17 ,953 40 ,955 1 0 , 2 4 4

Wing Group/S , psf 6.0 7.5 3.4 Emp. Grp/Semp, psf 3.5 3.9 2.3

Ultimate Load F a c t o r , g ' s 4.5 4.0 3.75


Wing, S 4 85 1 , 0 0 0 600 Horiz. T a i l . Sh 1 0 3 241 90

V e r t . T a i l , S, 90.2 170 49

Empenn. A r e a , S 1 9 3 411 139 emp


Table A l l . l a Group Weight Data f o r Flying Boats, I D I I P I ~ I = I I I P I I U ~ ~ ~ I I I I P P U I U U L . ~ I ~ I ~ ~ ~ ~ P ~ = = U ~ ~ ~ ~ =

Amphibious and F l o a t Airplanes ' P I P I I P a P u I I P u ~ P P ~ I a f P P P ~ P ~ P ~ u

a

TYPe A t t h e time of p r i n t i n g no d a t a were a v a i l a b l e

Number of engines: Weight Item, l b s


Nose Gear Main Gear


Engines P rope l l e r s Fuel System Propulsion System --------------- Power P lan t To ta l --------------- Avionics + Instrum. Surface Controls Hydraulic System Pneumatic System E l e c t r i c a l System Elec t ron ics Armament A i r Cond. System Anti- icing System Furnishings Auxil iary Gear --------------- Fixed Equipm't To ta l ---------------

Armament Provis ions Water

Max. Fuel Capacity Payload (Max.


Table A l l . l b Group Weight Data f o r Flying Boats, llPlPl~~PIPPIllmlPInnIII..ILIIttPPI=ILImII:ln=lI:n==

Amphibious and F l o a t Airplanes PPIPIIPIP=PPPE=IIPPIIIII~=E~==

F l i g h t Design Gross Weight, GW, l b s

StructurelGW Power Plant/GW Fixed Equipm'tIGW Empty WeightIGW

Wing GroupIGW Empenn. GroupIGW Fuselage GroupIGW Nacelle GroupIGW Land. Gear GroupIGW

Take-of f Gross Wht, WTO, l b s

Empty Weight, W E D I b s

Wing GroupIS, psf Emp. GrpIS emp* psf

A t t h e t i m e of p r i n t i n g no d a t a were a v a i l a b l e

U l t i m a t e Load Factor , g ' s

Surface Areas, f t 2

Wing, S Horiz. T a i l , Sh

V e r t . T a i l . S,

Empenn. Area, S emp

P a r t v Appendix A Page 185

T a b l e A12. l a Group Weight Data fo r Supersonic = = = = = = I P P ~ P ~ P I P P I I P P I ~ ~ E P ~ P ~ : P ~ P O I P I = ~ P = P ~ P ~ ~ ~ U

C r u i s e A i r p l a n e s IPPP===PPIPPPIDI

Supe r- AST-100 SSXJET c r u i s e r 8 * * ***

Number of e n g i n e s : 4 2 2 We igh t Item, l b s

Wing Group 85,914 3 , 5 9 9 3 ,962 Empennage Group 1 0 , 6 5 5 4 81 225 F u s e l a g e Group 52 ,410 3 , 4 9 4 2 ,195 Nacelle Group 1 6 , 8 0 3 505 700 Land. Gear Group 27 ,293 1 , 3 9 1 1 , 3 0 0

Nose Gear Main Gear .......................

S t r u c t u r e T o t a l 193 ,075 9 , 4 7 0 8 ,382 ....................... E n g i n e s 52 ,000 3 ,016 4 , 7 8 1 A i r I n d u c t . Sys t em i n c l . i n propulsion system F u e l S y s t e m 5 , 7 8 1 62 6 560 P r o p u l s i o n Sys t em 1 , 7 8 0 5 9 165 ....................... Power P l a n t T o t a l 59 ,561 3 , 7 0 1 6 ,756 ....................... A v i o n i c s + Instrum. S u r f a c e C o n t r o l s H y d r a u l i c Sys t em P n e u m a t i c S y s t e m E l e c t r i c a l S y s t e m E l e c t r o n i c s Armament A i r Cond. System**** A n t i - i c i n g Sys t em F u r n i s h i n g s A u x i l i a r y Gear

F i x e d E q u i p m l t T o t a l

5 , 050 4 4 5 357 i n c l . i n avionics + i n s t rum.

0 0 600 8 ,200 352

210 9 5 250

2 5 , 1 1 1 417 242 0 0 40 .......................

59 ,666 2 , 7 9 9 4 ,482

Mission F u e l Reqd. 327 ,493 18 ,674 1 2 , 5 2 3 P a y l o a d 61 ,028 725 5 ,000

*NASA TM X-7393 6, M=2.2 l a r g e p a s s e n g e r t ranspor t **NASA TM 74055, Ms2.2 e x e c u t i v e ( b u s i n e s s ) jet ***NASA TM 7 8811, M ~ 2 . 6 m i l i t a r y missile c a r r y i n g

s u p e r cruiser * * * * I n c l u d e s p ressur iza t ion system


T a b l e A12. lb Group Weight Data for S u p e r s o n i c = = = X P 5 = = ~ P P I = P = P P P = = = P P P P ~ l i P P ~ O 3 5 u I ~ a P ~ P P ~ ~ ~ ~ ~

C r u i s e A i r p l a n e s P===P=PSSPSP===P

Supe r- AST-100 SSXJET cruiser * * * * *

F l i g h t Design Gross Weight , GW, l b s 718 ,000 35 ,720 37 ,144

StructureIGW 0.269 0.265 0.226 Power PlantIGW 0.083 0 .104 0.182 F i x e d ~ q u i p r n ' t I G ~ 0.083 0 .078 0 .121 Empty Weight IGW 0.435 0.447 0 .528

Wing GroupIGW 0.120 0 .101 0.107 Empenn. GroupIGW 0 .015 0 .013 0.006 F u s e l a g e GroupIGW 0.073 0 .098 0.059 Nacelle GroupIGW 0.023 0.014 0.019 Land. Gear GroupIGW 0.038 0.039 0.035

Take- off Gross Wht, WTO, l b s 718 ,000 35 ,720 47 ,900

Empty Weight, WE, l b s 312 ,302 15 ,970 1 9 , 6 2 0

Wing GroupIS , p s f 8.6 3.7 10 .7 Emp. GrplSemp, p s f 11 .0 3.9 3 .1

U l t i m a t e Load F a c t o r , g 's 3.75 3 .75 6.0

S u r f a c e A r e a s , it2

Wing, S 9 ,969 965 371 Horiz. T a i l , Sh 579 62 0

V e r t . T a i l , SV 3 86 62 73


*NASA TM X-7393 6, M=2.2 large p a s s e n g e r t ransport **NASA TM 74055, M-2.2 executive ( b u s i n e s s ) jet ***NASA TM 7 8811, Ms2.6 m i l i t a ry missile c a r r y i n g

super cruiser e


Table A13. l a Group Weight Data f o r NASA X Airplanes X P = ~ P I = P P P P I I I P P ~ I I I E I ~ = = P X P ~ I = ~ = ~ P P P P ~ P = = ~ = ~ ~ ~ ~ ~ = ~

North TYPe Ryan American H i l l e r B e l l - X-13* X-IS** X-18*** XV-15 Number of engines: 1 1 2 2 Weight Item, l b s J e t Rocket TBP TBP

T i l t ro to r Wing Group 515 1,144 3,483 946 Empennage Group 146 1,267 928 259 Fuselage Group 415 3,806 4,694 1,589 Engine Sect ion 6 9 187 728 nac. 369 Land. Gear Group 300 3 89 1,289 524

Nose Gear Main Gear ...................................

S t r u c t u r e Tota l 1,445 6,793 11,122 3,687 ................................... E n g i n e ( ~ ) 2,766 6 80 5,460 1,052 F u e l System 100 1,354 62 3 226 Prope l l e r Ins t . 0 0 3,679 86 3 Propulsion System 227 14 8 548 222 Drive system 1,340 Prope l l e r Controls 0 0 2,023 **+*629 ................................... Power P l a n t Tota l 3,093 2,182 12,333 4,332 ................................... Avionics + Instrum. 41 172 141 231 Surface Controls 416 1,182 896 ****777 Hydraulic System 2 14 240 863 u t i l . 86 E l e c t r i c a l System 311 14 2 931 418 Elec t ron ics 29 175 63 Test Instrumentat ion 139 1,328 1,160 B a l l a s t 106 A i r Cond. System 10 192 119 Furnishings 19 9 446 81 3 43 4 Auxil iary Gear 0 11 0 10 ................................... Fixed Equipm't To ta l 1,359 3,888 3,707 3,341 ...................................

Liquid Nitrogen 150 engine o i l : 53 Max. Fuel Capacity 1,400 314 82 3 1,401 Payload crew and t e s t equipment only

*Delta conf igura t ion , took o f f from v e r t i c a l p o s i t i o n **Air launched by B-52 ***Turbo/propeller dr iven , wing incidence v a r i a b l e over

more than 90 degrees ****hydraulic system i n c l . i n r o t o r and su r face c t r l s


T a b l e A13. lb Group Weight Data fo r NASA X A i r p l a n e s I P P = P I ~ = P ~ P P = I P P I P P I P P I P ~ . c P P = = P P P ~ I P ~ P n 5 P ~ u n ~ P ~ n ~ u u ~

N o r t h .*++ TYPe Ryan Amer ican Hiller - B e l l

X-130 X-IS** X-18*** XV-15

F l i g h t D e s i g n Gross Weight , GW, l b s 7 ,000 1 3 , 5 9 2 33 ,000 1 3 , 2 2 6

S t r u c t u r e I G W 0.206 0 .500 0.337 0.279 Power P l a n t /GW 0.442 0 .161 0.374 0 .328 F i x e d EquipmPt/GW 0.194 0.286 0.112 0.253 Empty Weight /GW 0. 822 0.949 0. 826 0.859

Wing GroupIGW 0.074 0.084 0 .106 0.072 Empenn. GroupIGW 0.021 0.093 0 .028 0.020 F u s e l a g e GroupIGW 0.059 0.280 0.142 0 .120 E n g i n e Section/GW 0.010 0.014 0.022 0 .028 Land. Gear GroupIGW 0.043 0.029 0.039 0.040

Take- of f Gross Wht, WTOn I b s 7 ,149 1 3 , 5 9 2 3 3 , 0 0 0 1 3 , 2 2 6

Empty Weigh t , WE, l b s 5 ,755 1 2 , 9 0 1 27 ,272 1 1 , 3 6 0

Wing GroupIS , psf 2.7 10.9 6.6 5.6 Emp. G r p I S

emp' psf 2.3 10.0 2. 8 2.6

U l t i m a t e Load F a c t o r , g ' s 6.0 11.0


Wing, S 1 9 1 1 0 5 528 169 Horiz. T a i l , Sh o 52 201 50.3

V e r t . T a i l , Sv 62. 8 74.9 1 3 3 50.5


62. 8 127 3 34 1 0 1

*Delta configurat ion, t o o k off from vert ical pos i t i on **Air l a u n c h e d b y B-52 * * * T u r b o / p r o p e l l e r d r i v e n , wing i n c i d e n c e variable over

more t h a n 90 d e g r e e s * * * * T i l t r o t o r research airplane


Table A13 .2a Group Weight Data f o r NASA X Airp lanes = = = = = = = = ~ = 9 = ~ r ~ ~ ~ = ~ = = = a = = = ~ e = o = = = = i = t = ~ % 5 a = = = a ~ ~ = ~ a =

* a * *

TYPe B e l l B e l l Northrop B e l l X-2* X-5** YP-61*** XP-77

~ u n b e ; of eng ines : 1 1 2 1 Weight Item, l b s Rocket Jet P i s t o n / P i s t o n /

Prop. Prop. Wing Group 2,856 1,683 3,969 463 Empennage Group 445 19 8 62 9 5 9 Fuselage Group 4,108 1,064 1,557 218 Engine S e c t i o n 30 274 1,817 123

i n c l . booms Land. Gear Group 421 532 1, 803 344

Nose Gear 108 464 303 123 Main Gear ( s k i d s ) 313 6 8 1,500 221 ..................................

S t r u c t u r e T o t a l 8,281 3,751 9,775 1,207 .................................. E n g i n e ( ~ ) 607 2,223 4,974 738 Air Induc t ion System 3 1 Fuel System 89 8 108 914 91 Propuls ion System 13 77 1,081 101 P r o p e l l e r s 1,111 206 .................................. Power P l a n t T o t a l 1,518 2,439 8,080 1,136 .................................. Avionics + Instrum. 6 5 35 119 37 Sur face Cont ro l s 3 64 195 400 42 Hydraulic System 442 139 240 0 E l e c t r i c a l System 604 127 668 92 E l e c t r o n i c s 63 86 721 99 Ant i- ic ing System 100 T e s t In s t rumen ta t ion 708 155 Armament ( i n c l . guns and ammo 3,364 391 B a l l a s t 9 8 Air Cond. System 102 69 Furn ish ings 158 84 252 5 6 Auxi l i a ry Gear 0 90 352 15 .................................. Fixed Equipm't T o t a l 2,506 1,078 6,216 732 ..................................

Liquid Oxygen 7,180 Liquid Nitrogen 26 o i l : 23 ;il: 270 2 8 Max, Fuel Capaci ty 5,716 1,200 3,168 312 Payload crew and test equipment o n l y

*Air launched by B50 **Variable sweep wing ***Twin boom f i g h t e r ****Wood b u i l t l i gh twe igh t f i g h t e r


T a b l e A13.2b Group Weight Data f o r NASA X A i r p l a n e s P E l t l P P = P 3 t P I I P P P = P = P ~ P P P ~ ~ ~ = P 5 P P P O 3 1 ~ P ~ % ~ ~ ~ t P P ~ ~ ~ P P

* * * * TYPe B e l l B e l l N o r t h r o p B e l l

X-2* X-5** YP-61*** XP-77

F l i g h t D e s i g n Gross Weight, GW, l b s 25,627 8 ,737 27 ,813 3 , 6 3 2

S t r u c t u r e / G W 0.323 0 .429 0 .351 0.332 Power Plant/GW 0.059 0.279 0 .291 0.313 F i x e d E q u i p m l t /GW 0.098 0 .123 0.223 0.202 Empty Weight/GW 0.480 0. 832 0. 865 0.847

Wing Group/GW 0.111 0.193 0.143 0.127 Empenn. Group/GW 0.017 0.023 0.023 0 .016 F u s e l a g e Group/GW 0.160 0.122 0.056 0.060 E n g i n e Sect ionIGW 0.001 0 .031 0.065 0.034

i n c l . booms Land. Gear GrouplGW 0.016 0.061 0.065 0 .095

Take- of f Gross Whtn WTO, l b s 25,627 8 ,737 27 ,813 3 , 6 3 2

Empty Weigh t ,

W ~ ' l b s 1 2 , 3 0 5 7 , 2 6 8 24 ,071 3 ,075

Wing Group/S , psf 11 .0 9.6 6.0 4. 6 Emp. G r p / S

emp' p s f 4 .1 3.4 3.0 2.1

U l t i m a t e Load F a c t o r , g ' s no t a v a i l a b l e


Wing, S 260 175 6 64 100 Horiz. T a i l , Sh 49.5 31. 8 1 2 0 18.7

V e r t . T a i l , S, 5 8 25. 8 92 9.0


1 0 8 57.6 212 27.7

* A i r l a u n c h e d b y BSO * * V a r i a b l e sweep wing ***Twin boom f i g h t e r ****Wood b u i l t l i g h t w e i g h t f i g h t e r


T a b l e A13.3a Group Weight Data f o r NASA X A i r p l a n e s ~ P D I L I ~ P P P P ~ ~ ~ ~ P P P P ~ ~ ~ ~ P ~ I ~ ~ ~ I I . ~ P ~ P P P P ~ P ~ ~ ~ P = ~ ~ ~ ~ ~ ~

Mc Type Donne11 C o n v a i r NAA C o n v a i r

XF-88A XF-92A** YF-93A*** XFY-1 umber of engines: 2 1 1 1 Weight Item, l b s Jet Jet Jet TBP*

Wing Group Empennage Group F u s e l a g e Group E n g i n e S e c t ion Land. G e a r Group

Nose Gear Main Gear


E n g i n e ( s 1 F u e l S y s t e m P r o p e l l e r I n s t . P r o p u l s i o n Sys t em


A v i o n i c s + I n s t r u m . S u r f a c e C o n t r o l s H y d r a u l i c System E l e c t r i c a l S y s t e m E l e c t r o n i c s T e s t ins t rumenta t ion B a l l a s t Armament Guns or cannons A i r Cond. System F u r n i s h i n g s M i s c e l l a n o u s

F i x e d Equipm' t T o t a l

E n g i n e o i l Max. F u e l C a p a c i t y P a y l o a d Crew

2 , 0 4 8 1 , 6 9 4 2 ,640 1 , 8 7 7 472 590 444 62 3

3 ,267 2 ,149 2 ,850 1 , 0 8 4 29 0 44 1 5 7

9 86 7 64 1 , 3 8 2 46 6 193 1 5 5 254 g e a r s on 793 609 1 , 1 2 8 f o u r f i n s ...................................

6, 802 5 ,197 7 ,360 4 ,207 ...................................

75 2 3 1 8 94 4 ,404 4 ,440 1 0 , 5 9 3 1 , 8 3 9

829(ammo) N.A. 863(ammo) 0 230 230 230 200

* I n c l u d e s a f t e r b u r n e r s * * D e l t a wing c o n f i g u r a t i o n ***F-86 m o d i f i e d w i t h NACA f l u s h s i d e i n l e t s * * * * C o u n t e r- r o t a t i n g propeller d r i v e n t a i l s i t t e r (VTOL)


T a b l e A13.3b Group Weight Data f o r NASA X A i r p l a n e s P P I D S P I P I I P I I P P I P P I ~ ~ P = P P ~ ~ ~ ~ I : I C I C = P P I I C P = P I ~ ~ E D ~ ~ ~ ~ E ~

Mc Type Donne11 C o n v a i r NAA C o n v a i r

XF-88A XF-92A YF-93A XFY-1

F l i g h t D e s i g n Gross Weigh t , GW, l b s 2 0 , 0 9 8 1 1 , 6 0 0 21 ,846 1 4 , 2 5 0

Structure/GW 0.338 0 .448 0.337 0.295 Power Plant/GW 0.205 0.243 0.215 0 .388 F i x e d Equipmvt/GW 0.178 0.155 0.176 0.160 Empty Weight/GW 0.722 0.845 0 .728 0.844

Wing Group/GW 0.102 0.146 0 .121 0.132 Empenn. Group/GW 0.023 0 .051 0.020 0.044 F u s e l a g e GrouplGW 0.163 0.185 0.130 0 .076 E n g i n e S e c t i o n I G ~ 0 .001 0.000 0.002 0 .011 Land. Gear GroupIGW 0.049 0.066 0.064 0.033

Take- of f Gross Wht. WTO, l b s 2 0 , 0 9 8 1 1 , 6 0 0 2 7 , 7 8 8 1 5 , 1 8 5

Empty Weight, W E D l b s 1 4 , 5 1 2 9, 804 1 5 , 9 1 2 1 2 , 0 2 1

Wing Group/S , psf 5.9 4.0 8. 6 5. 3 Enrp. Grp/S .lap' P S ~ 4.3 7. 8 5.6 3.5

Ultimate Load F a c t o r , g ' s 11 .0 11 .0 11 .0 11.3


Wing, S 350 425 306 355 Horiz. T a i l , Sh N. A. 0 N. A. N. A.

V e r t . T a i l , S, N. A. 76.0 N. A. N. A.


1 0 9 76.0 79.6 1 7 6


Table A13.4a Group Weight Data f o r NASA X A i rp l anes ----------------=------------- ---------------- -------------=========teE'

* * * a

~ u m b e c of engines : Weight Item, l b s

Wing Group Empennage Group Fuselage Group Engine S e c t i o n Land. Gear Group

Nose Gear Main Gear

Lockheed XV-4A* 2

Jet

~ o c k h e e d Ryan XV-4B** XV-SA***

6 4 J e t Je t +

l i f t f a n 395 1,059 167 267

1,274 1,341 333 4 5 3 89 4 82 77 82

312 400

B e l l X- 22A 4

Turbo- s h a f t

7 89 131

1,324 610 432 94

338

S t r u c t u r e T o t a l 2,263 2,558 3,194 3,286 ................................... E n g i n e ( ~ ) (main) 872 758 913 1,191 Engine(s1 ( l i f t ) 0 1,500 1,855 Exhaust system 520 608 304 8 Fuel System 155 14 2 124 175 P r o p e l l e r I n s t . i nc lud ing d r i v e s : 2,147 Ducts and suppor t s : fwd 695, a f t 686, t o t a l : 1,3 81 Propuls ion System 101 88 80 246 ................................... Power P l a n t T o t a l 1,648 3,096 3,276 5,148 ................................... Avionics + Instrum. 73 133 73 121 Sur face Con t ro l s 4 86 65 5 440 1,2 56 Hydraulic System 6 2 116 115 162 E l e c t r i c a l System 376 394 19 6 376 E l e c t r o n i c s 29 3 5 40 237 T e s t i n s t rumen ta t ion 5 83 200 5 15 1,520 Auxi l i a ry g e a r 52 27 158 10 Air Cond. System 32 5 8 34 4 5 Furn ish ings 209 3 91 235 376 ----------------------------------- Fixed Equipmet T o t a l 1,902 2,009 1,806 4,103 ...................................

Engine o i l 20 62 12 22 Max. Fuel Capac i ty 1,147 3, 815 2,430 2,031 Crew 180 43 0 180 3 60 Payload 1,200

*Ejec to r t y p e VTOL * * L i f t eng ine t y p e VTOL L i f t f a n re- s e a r c h a i r p l a n e * * * * T i l t r o t o r r e s e a r c h a i r p l a n e


T a b l e A13.4b Group Weight Data f o r NASA X Airplanes PXtPP==IP=llSee=tP=PSPPPIPI=I==P=e:E.====~==~=PSiLiC=e===~==~=~~

Lockheed Lockheed Ryan B e l l XV-4A XV-4B XV-5A X-22A

F l i g h t D e s i g n Gross Weight, GW, l b s 7 , 2 0 0 1 2 , 0 0 0 9 , 2 0 0 1 4 , 7 0 0

S t r u c t u r e / G W 0.314 0.213 0.347 0 .224 Power Plant /GW 0.229 0 .258 0.356 0.350 F i x e d Equipmet/GW 0.264 0 .167 0.196 0.279 Empty Weight/GW 0. 807 0.639 0.900 0. 853

Wing Group/GW 0.049 0.033 0.115 0.054 Ernpenn. Group/GW 0.024 0.014 0.029 0.009 F u s e l a g e Group/GW 0.168 0.106 0.146 0.090 E n g i n e Sect ion/GW 0.034 0 .028 0.005 0 .041 Land. Gear Group/GW 0.040 0.032 0.052 0.029

Take- of f Gross Wht, WToI l b s 7 ,200 1 2 , 0 0 0 9 ,972 1 4 , 7 0 0

Empty Weigh t , WE, l b s 5, 813 7 , 6 6 3 8 ,276 1 2 , 5 3 7

Wing Group/S , p s f 3.4 3. 8 4.1 4.9 Emp. Grp/Semp, p s f 3.2 3 .1 2.6 1 .5

U l t i m a t e Load F a c t o r , g ' s 7.5 4.5 6.0 4. 5


Wing, S 1 0 4 104 260 1 6 0 Horiz. T a i l , Sh 26.4 26.4 52.9 20

V e r t . T a i l , S, 27.5 27.5 51.0 68.5


53.9 53.9 104 88. 5


APPENDIX B: DATA SOURCE FOR NON-DIMENSIONAL R A D I I OF f = S I P P = P P 6 ~ l P P I I P I I P P P P P P P ~ ~ . C P ~ P P P P ~ P P ~ ~ ~ ~ ~ P ~ E ~ ~ ~ P ~ ~

GYRATION FOR AIRPLANES = P P I I I P P t D S P P P I P n P P P P P

The purpose of t h i s appendix is t o p resen t tabula ted d a t a f o r non-dimensional r a d i i of gyra t ion of a i rp lanes . Actual moments of i n e r t i a can be est imated from t h e s e non-dimensional r a d i i of gyra t ion with t h e he lp of Equations 3 .7 through 3.8.

The t a b l e s a r e organized a s fol lows:

Table B l t Table B2: Table B3: Table B 4 : Table B5: Table B6: Table B7a: Table B7b: Table B7c: Table B8: Table B9a: Table B9br Table BlOa: Table Blob: Table Bloc: Table BlOd: Table B11 : Table B12:

Homebuilt p r o p e l l e r dr iven a i r p l a n e s S ing le engine p r o p e l l e r dr iven a i r p l a n e s Twin engine p r o p e l l e r dr iven a i r p l a n e s Agr icu l tu ra l a i r p l a n e s Business j e t s Regional tu rboprope l l e r dr iven a i r p l a n e s Jet t r a n s p o r t s Pis ton- propel ler d r iven t r a n s p o r t s Turbopropeller dr iven t r a n s p o r t s Mi l i t a ry t r a i n e r s F i g h t e r s ( J e t ) F igh te r s (P rope l l e r Bombers (Pis ton- Propel le r ) Bombers ( J e t ) M i l i t a r y p a t r o l a i r p l a n e s ( P r o p e l l e r ) Mi l i t a ry t r a n s p o r t s (P rope l l e r ) Flying boats Supersonic c r u i s e a i r p l a n e s

The d a t a i n a l l these table were der ived from manufacturers d a t a and/or from Ref.11.

P a r t v Appendix B Page 1 9 7

T a b l e B1 Non- dimens iona l R a d i i of Gyration f o r Homebui l t P r o p e l l e r P ~ I P P P a P l ~ ~ P P t P P P ~ P ~ = ~ ~ ~ P P P D . ~ P P ~ P P ~ ~ ~ ~ ~ ~ ~ ~ = n = ~ ~ ~ a ~ m ~ ~ ~ E ~ n ~ ~ n ~ u w ~ = ~

Driven A i r p l a n e s P P m P l P P P P P I I P P a P

- - - A i r p l a n e Type GW Wing T o t a l e = Rx 5. Rz Number o f

Span, L e n g t h , ( b + L ) / 2 , e n g i n e s and l b s b , it L, f t f t d i s p o s i t i o n

At t h e time of p r i n t i n g , no d a t a were a v a i l a b l e for t h i s type airplane

T a b l e B2 Non- dimensional R a d i i of Gyration f o r S i n g l e E n g i n e o ~ ~ m % ~ ~ ~ ~ ~ ~ m ~ ~ ~ ~ ~ ~ ~ n ~ ~ ~ ~ n ~ a ~ = i i i u i e : ~ ~ ~ a ~ ~ ~ ~ a m ~ n a m m ~ ~ ~ ~ ~ ~ n a ~ a a ~

P r o p e l l e r Driven A i r p l a n e s = ~ I D P m P P I w ~ P P P ~ I I P P P ~ P P P P 0 5

- - A i r p l a n e Type GW Wing T o t a l e - Rx R

Span , L e n g t h , (b+L) / 2 , Y l b s b , f t L, f t it

Beech N-35. 3 , 1 2 5 32.8 25.1 29.0 0 .248 0.338 C e s s n a 150M** 1 , 1 2 7 33.5 21.5 27.5 0.254 0.405 C e s s n a 172M** 1 , 4 7 7 36.2 26.5 31.4 0.242 0.386 C e s s n a 177A** 1 , 7 6 1 35.6 27.0 31.3 0 .212 0.362 C e s s n a R182* 1 , 8 8 5 36.2 28.0 32.1 0.342 0.397 C e s s n a 210K*** 2 ,700 36 .8 28.3 32.6 0.222 0 .356

*at WTO * * a t WOE ***atWOE p lus 2 5 percent f u e l

- RZ

Number of e n g i n e s a n d d i s p o s i t i o n

0.393 1 i n fuse l . 0 .418 1 i n fuse l . 0 .403 1 i n f u s e l . 0.394 1 i n f u s e l . 0.393 1 i n f u s e l . 0 .379 1 i n fusel .

Note: one p i l o t i n c l u d e d i n a l l d a t a

Table B3 Non-dimensional Radi i of Gyration f o r Twin Engine P P I P P 6 P P P P P P O P I P P m P P n L : P I ~ P P P ~ . c P P I P 5 P P P = ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ P ~ ~ ~ ~ ~ ~

Prope l l e r Driven Airplanes P P P I O P P I P I P I P P P P P I P I P P P L i = ~

- - Airplane Type GW Wing Tota l e n Rx Rz Number of

Span, Length, (b+L) 12, engines and l b s b, f t L, f t f t d i s p o s i t i o n

Beech 55 Beech 95 Beech D-50 Beech D18S Cessna 402* Cessna 402 Cessna 404. Cessna 404 Cessna 441* Cessna 441

0.260 0.329 0.399 2 on wing 0.251 0.327 0.391 2 on wing 0.240 0.313 0.384 2 on wing 0.232 0.360 0.396 2 on wing 0.414 0.278 0.502 2 on wing 0.373 0.269 0.461 2 on wing 0.324 0.318 0.446 2 on wing 0.340 0.284 0.445 2 on wing 0.285 0.345 0.429 2 on wing 0.256 0.212 0.336 2 on wing

Table B4 Non-dimensional Radii of Gyration f o r Agr icu l tu ra l Airplanes l ~ l l ~ P P P P I P P L P P P I P P I P l P I P P ~ ~ ~ ~ i P P J ~ ~ P ~ ~ ~ ~ ~ ~ ~ U P ~ P ~ ~ ~ ~ P ~ ~ U ~ 5 ~ ~ ~ ~ ~ ~ I ~ ~ ~ ~ ~

- - - Airplane Type GW Wing To ta l e - St 5. Rz Number of

Span, Length, (b+L) 12, engines and l b s b, f t L, f t f t d i s p o s i t i o n

'I

A t t h e t i m e of p r i n t i n g , no d a t a were a v a i l a b l e f o r t h i s type of a i r p l a n e

T a b l e B5 Non- dimens iona l R a d i i of Gyration for B u s i n e s s Je ts u - ~ u u = = m = = ~ P n = s = ~ u = = = = u u m u = - = = P n n u = n = P u I c u = - n a ~ n ~ = ~ = = m n u = = n = =

- - A i r p l a n e Type GW Wing T o t a l e - Rx % Rz ~ u m b d r of

Span, L e n g t h , ( b + L ) / 2 , engines a n d l b s b, f t L, f t f t d i s p o s i t i o n

Morane/S 760 7 ,066 33.3 32.9 33 .1 0.374 0 .328 0.486 2 i n W/F Lockh. Jetstar 3 9 , 2 8 8 53.7 58 .8 56.3 0.370 0 .356 0.503 4 on f u s e l . C e s s n a 500. 6 ,505 47.1 43.5 45.3 0.236 0.384 0 .430 2 on f u s e l . C e s s n a 500** 1 2 , 0 0 0 47.1 4 3 - 5 45.3 0 .306 0.303 0.423 2 on f u s e l . C e s s n a 550. 7 ,036 51.7 4 7 - 2 49.5 0.243 0 .400 0.447 2 on f u s e l . C e s s n a 550** 1 3 , 5 0 0 51.7 47.2 49.5 0.293 0.312 0.420 2 on fu se l .

T a b l e B6 Non- dimens iona l R a d i i of Gyration for R e g i o n a l T u r b o p r o p e l l e r P P ~ ~ I ~ I I E ~ E I ~ I P P ~ ~ ~ ~ P ~ ~ ~ ~ ~ I ~ ~ ~ ~ ~ ~ ~ E ~ E ~ P ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ E ~ E ~ ~

Driven A i r p l a n e s uuu-==n=u=a-==Pu

- - - A i r p l a n e Type GW Wing T o t a l e = % % Rz Number of

Span, L e n g t h , (b+L) / 2 , engines a n d l b s b, f t L, f t f t d i s p o s i t i o n

F o k k e r F-27A 38 ,500 95.2 77.2 86.2 0.235 0 .363 0 ,416 2 on wing DHC6 Twin O t t e r 1 2 , 5 0 0 65.0 51.8 5 8.4 0.203 0.326 0.350 2 on wing

Table B7a Non-dimensional R a d i i of Gyrat ion for Jet Transpor t s

Airplane Type GW

l b s

Conva i r 8 80 185,000 Convair 8 80 191,500 Convair 990 240,000 Convair 990 245,000 Boeing 727-100 165,000 Boeing 727-100. 89,000 Boeing 727-200 180,000 Boeing 727-200* 100,000 Boeing 737-200 113,000 Boeing 737-200* 62,000 Boeing 747-100B 800,000 Boeing 747-100B* 350,000 McDD DC9-10 74,000 MCDD DC8 210,000

Wing Span, b, f t

- - T o t a l e - Length, ( b + ~ ) 12,

Rx f L, f t it

Number of engines and d i s p o s i t i o n

4 on wing 4 on wing 4 on wing 4 on wing 3 on f u s e l . 3 on f u s e l . 3 on f u s e l . 3 on f u s e l . 2 on wing 2 on wing 4 on wing 4 on wing 2 on f u s e l . 4 on wing

T a b l e B7b Non- dimens iona l R a d i i of Gyrat ion fo r P i s t o n- P r o p e l l e r Driven T r a n s p o r t s - - ~ P P ~ P I P S P = = P ~ = ~ ~ P ~ ~ O P P L ~ P P L ~ P ~ = P P P = P ~ P ~ P ~ E U ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ = ~ ~ = ~ ~ ~ ~ ~ U ~ ~ P ~ ~ 5 = X = l = P S I - -

- - - A i r p l a n e Type GW Wing T o t a l e = Rx % Rz Number o f

Span , L e n g t h , ( b + L ) / 2 , e n g i n e s a n d l b s b , f t L, f t f t d i s p o s i t i o n

Lockheed L-749A 107 ,000 123.0 95.2 109 .1 Lockheed L-1049 120 ,000 123.0 113.6 118.3 Lockheed L-1649 1 4 6 , 5 0 0 150.0 116.2 133 .1 Doug la s DC-4 60 ,360 138 .3 97.6 118.0 Doug la s DC-6 97 ,200 117.5 100.5 109.0 A i r s p e e d Ambass. 49 ,500 115.0 80.4 97.7 Martin 404 45 ,000 93.3 74.6 84.2 C o n v a i r T-240 41 ,800 91.7 74.7 83.2 C o n v a i r T-340 44 ,500 105.7 79.2 92.4 Beech Twin Quad 20 ,000 70.0 52.7 61.4

L 0.300 0 .298 0.426 4 on wing 0 .316 0 .336 0 .448 4 on wing 0 .371 0 .278 0.473 4 on wing 0.250 0.320 0 .388 4 on wing 0.322 0.324 0.456 4 on wing 0 .278 0 .314 0.400 2 on wing 0.272 0 .378 0.444 2 on wing 0 .286 0 .351 0.443 2 on wing 0.308 0 .345 0.457 2 on wing 0 .225 0.303 0.346 4 i n wing

T a b l e B7c Non-d imens iona l R a d i i of Gyration for T u r b o- P r o p e l l e r D r i v e n T r a n s p o r t s

- - - A i r p l a n e Type GW Wing T o t a l e E Rx 5. Rz Number of

Span , L e n g t h , (b+L) / 2 , engines a n d l b s b, f t L, f t f t d i s p o s i t i o n

B r i s t o l 175 (k)* 103 ,000 130.0 110.0 120 .0 0.317 0.356 0 .455 4 on wing B r i s t o l 167 (1 )**187 ,000 230.0 177 .0 203.5 0.330 0 .356 0 .478 4 on wing Lockh. E l e c t r a 116 ,000 99.0 104.7 101 .9 0.394 0 .341 0.497 4 on wing * B r i t a n n i a **Brabazon

T a b l e B8 Non- dimens iona l R a d i i of Gyration f o r M i l i t a r y T r a i n e r s P P P I I P P P I P I P P P 1 I ~ P I P U U ~ P P P P E P I P P E ~ ~ P P E P P ~ ~ ~ ~ ~ E ~ ~ ~ ~ ~ ~ I ~ ~ ~ ~ ~ ~ E ~ 6 ~ ~

A i r p l a n e Type GW Number of

Wing T o t a l e =

Span , L e n g t h , ( b + ~ ) /2 l b s b, f t L, f t f t

e n g i n e s and d i s p o s i t i o n

C e s s n a T-37A 6 ,300 38.4 30.0 34.2 0 .220 0.142 0 .245 2 i n f u s e l .

Table B9a Non-dimensional R a d i i of Gyra t i on f o r F i g h t e r s ( Je t ) P P U P P ~ I ~ P I P ~ ~ C P ~ I P O P P P P I ~ P ~ ~ ~ U ~ P ~ ~ ~ ~ ~ P ~ ~ ~ P ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ P P ~ ~

- - - Airp l ane Type GW Wing T o t a l e -

Span, Length, (b+L) / 2 , Rx % Rz

l b s b , f t L, f t f t

McD F2H-1 McD F3H-2N McD F-1O1A VS A t t a c k e r DH Vampire 20 G1. Meteor I1 Lockheed F- 80A Lockheed F-94B Lockheed F-104G NAA F-86A NAA FJ-3 NAA F-100D Vought XF8U-1 Vought F8U-3 GD XF-91 GD TF-102A GD F-106B Northrop F-89D Republ ic RE'- 84F Republ ic F-105D Grummn F9F-8 Grumman XFlOF-1 Grumman F11F-1

'I

Number of eng ines and d i s p o s i t i o n

2 i n W/F 1 i n f u s e l . 2 i n f u s e l . 1 i n f u s e l . 1 i n f u s e l . 2 i n wing 1 i n f u s e l . 1 i n f u s e l . 1 i n f u s e l . 1 i n f u s e l . 1 i n f u s e l . 1 i n f u s e l . 1 i n f u s e l . 1 i n f u s e l . 1 i n f u s e l . 1 i n f u s e l . 1 i n f u s e l . 2 i n f u s e l . 1 i n f u s e l . 1 i n f u s e l . 1 i n f u s e l . 1 i n f u s e l . 1 i n f u s e l .

T a b l e B9b Non- dimens iona l R a d i i of Gyration fo r F i g h t e r s ( P r o p e l l e r ) ~ P ~ P P ~ P ~ P I P P ~ P P ~ P I I ~ ~ : P P P ~ ~ P P I ~ ~ P P ~ ~ ~ P ~ ~ P ~ ~ ~ ~ ~ ~ = ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ P ~ ~ P ~ ~ ~ I ~

- - - A i r p l a n e Type GW Wing T o t a l e - Rx % Rz Number of

S p a n , L e n g t h , (b+L) / 2 , engines a n d l b s b , f t L, f t f t d i s p d s i t i o n

~ r e w s t e r B u f f a l o 5 , 0 6 6 35.0 26.0 30.5 S e v e r s k y P35 5 , 7 8 8 36.0 26 .8 31.4 VS S p i t f ire- I 6 ,250 36 .8 29.9 33.4 BP Defiant 6 ,410 39.4 35.0 37.2 C u r t i s s P36 6 ,825 37.3 31.7 34.5 B e l l P39 7 , 5 3 3 34.0 30.0 32.0 Grurnman F6F 1 0 , 5 6 0 42 .8 33.5 38.2 Hawker Typhoon 1 1 , 0 1 7 41.5 31.7 36.6 R e p u b l i c P47 1 2 , 5 0 0 40.7 36.0 38.4 Vought F4U 1 2 , 8 5 0 41.0 34.5 37.8 B 1 . F i r e b r t d - I 1 1 1 3 , 6 6 0 49 .8 38.2 44.0 W e s t l a n d Welk in 1 8 , 3 4 0 70.0 42.0 56.0 B r i s t o l B e a u f t r 22 ,635 57.8 42.5 50.2 B r i s t o l B r i g a n d 39 ,000 71.7 46.4 59.1

0 .208 0 .358 0.374 1 i n f u s e l . 0 .198 0.367 0 .360 1 i n f u s e l . 0.240 0.334 0 .384 1 i n f u s e l . 0.234 0.360 0.404 1 i n fuse l . 0.172 0.356 0 .370 1 i n f u s e l . 0 .276 0 .340 0.425 1 i n f u se l . 0.242 0.346 0.404 1 i n f u s e l . 0.277 0.300 0.394 1 i n fusel . 0.296 0.322 0 .428 1 i n f u s e l . 0 .268 0 .360 0.420 1 i n f u s e l . 0.250 0 .300 0.397 1 i n f u s e l . 0 .270 0.304 0 .408 2 i n wing 0.330 0.299 0.447 2 i n wing 0.299 0 .338 0 .438 2 i n wing

T a b l e BlOa Non- dimens iona l R a d i i of Gyra t ion f o r Bombers ( P i s t o n - P r o p e l l e r ) ~ ~ ~ ~ E ~ P P P P I ~ L P I I ~ P I ~ P ~ ~ P ~ P P P ~ ~ P P ~ ~ P P ~ P ~ . ~ ~ P ~ E ~ ~ ~ ~ = ~ ~ ~ E ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~

- - - A i r p l a n e Type GW Wing T o t a l e n Rx 5. Rz Number of

Span , L e n g t h , (b+L) / 2 , engines and l b s b , f t L, f t f t d i s p o s i t i o n

Martin B-26 26 ,600 65.0 57.6 61.3 0.270 0.320 0.410 2 on wing HP H a l i f a x 55 ,000 99.0 71.6 85.3 0.346 0.306 0.395 4 on wing S h o r t s S t i r l i n g 64 ,000 99.0 87.3 93.2 0.360 0 .330 0 .488 4 on wing Boe ing B-29 105 ,000 141 .0 99.0 120.0 0.316 0 .320 0 .376 4 on wing Boeing B-50 120 ,000 141.0 99.0 120.0 0.304 0.332 0.450 4 on wing GD B-36 357 ,500 230.0 162 .0 196 .0 0.316 0 .262 0 .428 6 i n w i n g

T a b l e B l o b Non- dimens iona l R a d i i of G y r a t i o n for Bombers (Je t ) ~ U ~ U ~ ~ ~ P P P ~ P ~ P I P ~ ~ ~ ~ ~ ~ P ~ U P ~ ~ P ~ P . L ~ ~ P ~ P ~ P P ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ P ~ ~ ~ ~ ~ ~ ~ ~ ~

'd & - - 1 rt

A i r p l a n e Type GW Wing T o t a l e = Span , L e n g t h , (b+L) 1 2 ,

Rx

C 5.

l b s b , f t L, f t f t

Martin XB-51 53 ,785 53.0 81.0 67.0 0.194 0.404 Martin B57A 48 ,554 64.0 66.0 65.0 0.312 0 , 2 7 8 Boe ing XB-47 1 2 5 , 0 0 0 116 .0 107.0 111.5 0.346 0.320 Boe ing B-52A 390 ,000 185.0 156.5 170 .8 0.346 0.306 Northrop RB-49A 213 ,500 172 .0 53.0* 112.5 0.316 0 .316 NAA B-45A 82,600 89.0 75.3 82.2 0.325 0.290 NAA B-45C 82,600 89.0 75.3 82.2 0.340 0 .299

* F l y i n g wing

Number of e n g i n e s a n d d i s p o s i t i o n

3 i n l o n f u s . 2 4 i n i n wing wing

8 on wing 6 i n wing 4 on wing 4 on wing

T a b l e B l o c Non-dimensional R a d i i of Gyration for Military P a t r o l A i r p l a n e s = u ~ n = = ~ = = = ~ = = = = = = = = = n = ~ ~ = ~ = = u ~ = m = = = = = n = ~ u = = n = ~ ~ n = n ~ m ~ ~ a ~ a = n ~ ~ = ~ = ~ = u = ~ ~ = ~ a ~

( P r o p e l l e r D r i v e n ) ======1=1===1=111=

- - - A i r p l a n e Type GW Wing T o t a l e = Rx % Rz

Number of Span, L e n g t h , ( b + ~ ) 1 2 , engines a n d

l b s b , f t L, f t f t d i s p o s i t i o n

D-~VW L o c k h e e d ~ 2 V - 4 67 ,500 100.0 81.6 90.8 0.368 0.300 0.484 2 on wing ~ o c k h e b d P~V-7* 67,500 100.0 91.7 95.9 0.372 0.266 0.462 4 on wing Grumman S2F-3 26 ,147 69.7 43.5 56.6 0 .240 0.347 0.387 2 on wing Grumman WZF-1 41 ,549 80.6 56.3 68.4 0.235 0.366 0.387 2 on wing

Lockheed P3V-1 127 ,200 99.7 116 .8 108.3 0.357 0.249 0 .421 4 on wing

Table BlOd Non- dimensional R a d i i of G y r a t i o n f o r M i l i t a r y T r a n s p o r t s ~ P ~ P = ~ I I = P ~ ~ = = = ~ P P I P ~ ~ ~ ~ : = ~ P P P E ~ ~ = = ~ = Z I ~ ~ P ~ = ~ ~ = ~ ~ ~ ~ ~ ~ ~ ~ E ~ ~ ~ ~ ~ ~ ~ ~ ~ = ~ ~ = ~

( P r o p e l l e r Driven) = P P P P P I P P P I P a = I P D =

- - - A i r p l a n e Type GW Wing T o t a l e = Rx 5. Rz Number of

Span, L e n g t h , (b+L) 1 2 , e n g i n e s a n d l b s b , f t L, it f t d i s p o s i t i o n

P i s t o n Enz -~JJe r Driven - Doug la s C-54 61 ,840 117.5 94.0 105. 8 0.286 0.294 0.406 4 on wing F a i r c h i l d C-119B 64,000 109.3 88.5 98.9 0.287 0.282 0.390 2 On wing i n wing Boe ing C-97 1 2 8 , 3 4 0 141.2 110.3 125 .8 0.276 0 .325 0.424 4 on wing GD XC-99 265 ,000 230.0 182.5 206.3 0.276 0.346 0.432 6 on wing - r D r m

- Lockheed C-130B 1 3 5 , 0 0 0 132 .6 97. 8 115.2 0.363 0.319 0.489 4 on wing Lockheed C-130E 155 ,000 132 .6 97 .8 115.2 0.375 0 .290 0.486 4 on wing

*Has two jet e n g i n e s o u t b o a r d o f t h e p i s t o n e n g i n e s

T a b l e B11 Non- dimens iona l R a d i i of Gyra t ion f o r F l y i n g B o a t s I = P D P = = P = P P P = P = P P P P I P P = P i 5 : P ~ P P ~ = = P ~ P I P W O O D i = = ~ = = U S = P ~ ~ U ~ E ~ ~ = = = ~

- - A i r p l a n e Type GW Wing T o t a l e w Iix % Rz

Number of S p a n , L e n g t h , (b+L) 1 2 , e n g i n e s a n d

l b s b , f t L, f t f t d i s p o s i t i o n

XBTM-1 20 ,009 50.0 41.2 45.6 0 .230 0 .340 0.380 1 i n fuse l . XP4M-1 80,000 114.0 84.0 99.0 0 .248 0 .320 0 .414 2 on wing VS S e a g u l l 1 4 , 2 3 0 52.5 44.0 48. 3 0.297 0.364 0 .402 1 i n WF

T a b l e B12 Non- dimensional R a d i i of Gyration for S u p e r s o n i c C r u i s e A i r p l a n e s = P ~ P I ~ I P I ~ ~ P I P = P P I I P S ~ ~ P P ~ = ~ ~ P E ~ ~ ~ P ~ ~ E P P P ~ ~ ~ ~ ~ ~ ~ ~ ~ U ~ ~ = ~ ~ ~ ~ = ~ ~ ~ ~ ~ U ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~

A i r p l a n e Type GW Wing T o t a l e = Rx RY Rz Number of

Span, L e n g t h , ( b + ~ ) 12 , e n g i n e s a n d l b s b , f t L, f t f t d i s p o s i t i o n

NAA A3J-1 44 ,305 53.0 72.5 62. 8 0 .240 0.372 0.472 2 i n f u s e l .

1 2 . INDEX =========

Aerodynamic c e n t e r Af te rburner (weight ) A i r cond i t i on ing system (weight) A i r i nduc t ion system (weight) Ant i- ic ing system (weight Area Armament (weight Aux i l i a ry gea r (weight ) Aux i l i a ry power u n i t (weight Avionics (weight)

Baggage handl ing equipment (weight) B a l l a s t (weight) Braced wing

C a l i b r a t i o n (of weight equa t ion ) 2 9 Canard (weight) 5 C a n t i l e v e r wing 6 7 Cargo handl ing equipment (weight 1 1 0 , 9 7 , 5 Cessna method 1 0 7 , 1 0 1 , 9 8 , 9 3 , 9 0 , 8 4 , 8 0 , 7 8 ~ 7 5 , 7 1 , 6 7 , 2 6 Class I weight and balance: s e e P a r t I1 C l a s s I1 weight and ba lance 1 1 7 C l a s s I1 i n e r t i a s 1 2 1 C l a s s I weight e s t i m a t i o n method 4,3 ,2 ,1 . C l a s s I1 weight e s t i m a t i o n method 4 6 , 2 7 , 2 5 , 2 , 1 Component weights 2 Crew weight 2 7 , 2 6 , 4

De-icing system (weight) Design c r u i s e speed Design d i v e speed Design load f a c t o r : see load f a c t o r Design maneuvering speed

E l e c t r i c a l system (weight) E l e c t r o n i c s (weight) Empennage (weight Empty weight Engine (weight)

FAR 2 3 FAR 2 5 Fixed equipment c.g. l o c a t i o n Fixed equipment weight F l i g h t c o n t r o l system (weight) F l i g h t des ign g r o s s weight F l i g h t tes t in s t rumen ta t ion

P a r t v Index Page 2 0 7

Fuel f r a c t i o n s F u e l system (weight ) Fuel t r a n s f e r system Furn ish ings (weight) ~ u s e l i i g e (weight

GD method 1 0 0 , 9 9 , 9 1 , 8 9 , 8 8 , 8 7 , 8 1 , 7 9 , 7 7 , 7 6 , 7 3 , 7 0 , 6 9 , 2 6 110,109,108,106,105,103,102

Gross take- off weight 2 6 , 4 , 3 Guns (weight 1 1 1 , 9 7 , 6 Gust l oad f a c t o r l i n e 3 8 ,34

Hor izonta l t a i l (weight) Hydraulic system (weight )

I n e r t i a (moment and /o r p roduc t ) In s t rumen ta t ion (weight)

Landing g e a r (weight Launchers (weight) Load f a c t o r , l i m i t Load f a c t o r , u l t i m a t e (= 1 . 5 x l i m i t ) Locat ing component c.g. 's

Main g e a r (weight) Maximum d i v e speed Maximum g u s t i n t e n s i t y speed Maximum l e v e l speed Maximum Mach number a t s e a l e v e l Maximum ze ro f u e l weight Mission f u e l weight Moving component c .g . ' s Moving t h e wing

Nacelle (weight) Negative s t a l l speed Norm1 f o r c e c o e f f i c i e n t Nose g e a r (weight)

O i l system and o i l c o o l e r Ope ra t iona l i t ems (weight) Out r igger g e a r (weight) Oxygen system (weight)

P a i n t (weight) Payload weight Pneumatic system (weight) Powerplant c.g. l o c a t i o n Powe r p l a n t weight P r e s s u r i z a t i o n system (weight )

P a r t V Index Page 208

Prope l l e r (weight) Propulsion system (weight)

Radius of qvra t ion Radius of gkra t ion (non-dimensional) d a t a Ramp ( i n l e t )

S t a l l speed S t r u c t u r a l arrangement S t r u c t u r a l component c. g. l o c a t i o n S t r u c t u r e weight S t r u t t e d wing ( s t r u t braced wing) Sweep angle

1 8 , 1 7 , 2 Appendix B

8 8

T a i l boom (weight) 5 T a i l gear (weight) 5 Taper r a t i o 25 Thickness r a t i o 2 5 Thrust reverser (weight 1 95,s Torenbeek method 9 1 , 9 0 . 8 8 . 8 7 , 8 4 n 8 1 n 7 9 , 7 7 , 7 4 , 7 3 . 6 9 . 6 8 . 2 6

109 ,108 ,106 ,105 ,104 , 103,102,99, 95, 92 Trapped f u e l and o i l weight 2 6 , 4

USAF method

V e r t i c a l t a i l (weight) V-n diagram

Water i n j e c t i o n system 96 Weapons provis ions (weight) 1 1 1 , 9 7 , 6 Weight d a t a Appendix A

a g r i c u l t u r a l s 1 3 5 business jets 1 4 0 , 1 3 8 f i g h t e r s 172 ,170 ,168 ,166 ,164 f l y i n g boats 1 8 4 homebuilts 1 2 6 j e t t r a n s p o r t s 158 ,156 ,154 ,152 ,150 m i l i t a r y j e t t r a n s p o r t s 174 m i l i t a r y p a t r o l a i r p l a n e s 182 m i l i t a r y p i s ton lp rop t r a n s p o r t s 1 8 0 , 1 7 8 m i l i t a r y t r a i n e r s 162 m i l i t a r y turboprop t r a n s p o r t s 1 7 6 NASA experimental a i r p l a n e s 1 9 4 , 1 9 2 , 1 9 0 , 1 8 8 r eg iona l s 148 ,146 ,144 ,142 s i n g l e engine props 1 3 0 , 1 2 8 supersonic c r u i s e a i r p l a n e s 1 8 6 turboprop t r a n s p o r t s 1 6 0 twins 1 3 4 , 1 3 2

Weight f r a c t i o n ( s 1 Appendix A,7 ,6 ,4 ,2 Wing (weight 67 ,s

P a r t V Index Page 209

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