Initial aircraft weight prediction

fÊCHNISCHE HCGT^SCH

THE COLLEGE OF AERONAUTICS

CRANFIELD

**^

INITIAL AIRCRAFT WÉiGHT PRED?CTiON

by

T H E C O L L E G E O P A E R O N A U T I 0 S

C R A N F I E L D

I n i t i a l A i r c r a f b ¥ei,p:ht Frediction^o

„•by-D, HovTe, D.C.Ae.

Although much has recently been published oonceming methods for the weight prediction of individual structural components, it is some considerable time since a revle\7 of the subject as a whole has appeared. This note is an attempt to remedy the situation, the scope of the work not being limited to aircraft structures but including also equipment and systems,

Various stages in the design process are considei-ed, commencing yrith simple formulae suitable for use in the operations system

stage, passing through the moi-e detailed project stage and finally recommending methods suitable for weight prediction as the design becomes established,

In some instances the suggested formulae Hre vrell kncwn^ but an effort has been made to utilise the latest available

information in bringing techniques up to date and where possible to fill in gaps, partic\ilarly with respect to equipment and systems.

COKTEIWS

Notation.

1,0 Introduction.

2.0 S t r u c t u r e ,

2.2 Fuselage

2.3 T a i l Unit

2.4 Nacelles and Pods

2.5 T a i l Booms

2.6 Undercarriage

2.7 Floats

2.8 Flaps

3,0 Power Plant,

4.0 Systems.

4.1 Flying Controls

4.2 Hydraulics, E l e c t r i c s and Pneumatics

4 . 3 De-icing

4.4 Pressurisation and Air Conditioning

4.5 Fuel System

5.0 Equipment and miscellaneous items,

5.1 Seats and Furnishings,

5.2 Intruments

5.3 Radio and Radar

5.4 Fire Precautions

5.5 Oxygen Eqioipment

5.6 Paint

References,

3

-Notation.

A Aspect Ratio. b Wing Span (feet),

B Maximum Fuselage ¥/idth ( f e e t ) .

C Terms in wing vraight formula, see E q . ( l ) . '

d Tyre diameter ( i n c h e s ) .

D Term i n formula for \7ing v/eight, see Eq„(2) ,

Q F Ratio of Ultimate Tensile Strength of m n g naterial at 15 C

to that at tenxperature corrx3sponding to M.

f Drag factor in brake weight formula, see paragraph 2.6. g Location of inertia axis aft of leading edge as ratio of

chord but not less than 0.4. h Maximum fuselage depth (feet) ,

H ^"V- for A < 3 or cos '^ J\- for A > 3 0.9(1+^)

k Taper Ratio c o r r e c t i o n , see F i g . 1 ,

n Distance between the ground l i n e vri-th undercarriage leg

^ extended and mean point a t v/hich load i s t r a n s f e r r e d t o

the airframe, p a r a l l e l to the l e g ( i n c h e s ) ,

L Ratio of the Shear Modulus of the wing m a t e r i a l a t 15 C

to t h a t a t temperature corresponding to M,

L„ T a i l Arm ( f e e t ) .

M Mach Number corresponding t o V_,,

N Maximum factored normal a c c e l e r a t i o n .

n NiJmber of b r a k e s .

p Tyre pres store ( l b s / s q , i n ) ,

P Max. factored r e s i i l t a n t r e a c t i o n , normo-l t o \mdercarriage leg ( l b s ) .

Q Factor i n ¥ing formula, see ^paragraph 2.1 ,

r Relief Factor = 1 - R/\7,

R F i r s t Moment of Relief Loads divided by 0,2b ( l b s ) ,

s Number of engines. ^

S Wing Area ( s q . f t ) ,

Sp G r o s s fuselage surface a r e a ( s q . f t ) , Sj, G r o s s h o r i z o n t a l t a i l a r e a ( s q . f t ) ,

Notation (continued).

Sy. Gross vertical t a i l area ( s q . f t ) , t Flight time in hours,

T Engine thrust ( l b s ) .

t/c) Thickness chord r a t i o of wing at root, u Tifheel rim diameter (inches).

U Maximum dynamic wheel reaction. V Tank Volume (gallons).

V^ Design diving speed ( f . p . s , ) . V_ Stalling Speed ( f . p . s . ) , w Tyre width (inches),

¥ Take off weight (excluding overload or boosts)

\1 Brake Weight ( l b s ) .

W„ ï'uselage YJeight ( l b s ) ,

W Weight of Horisontal Tail ( l b s ) . W, Lending Weight (lbs) ,

Yiw,, Weight of Main U n d e r c a r r i a g e ( l b s ) , W|^- Y/eight of Nose Undercairiage ( l b s ) . W^ Tyre Vfeight (lbs)

W™, Weight of Tail Undercarriage ( l b s ) . W„ Weight of Vertical Tail ( l b s ) , VL, Y/ing Yfcight (lbs) .

A

a I Terms in fuselage weight formula, see Eq,(3).

Y Svreepback of Structure,

y\^ Leading edge ^cepback,

/ Sweepback of 0,25 chord,

5

-1, Introduction.

It is not necessary to eiiiphasise the importance of correct weight estimation in the early design stages of on aircraft. Althoiigh a nijmbor of contributions to the subject have been forthcoming in recent years, those ho.ve been mainly applicable to a particular structural component and it is some considerable time since .the overall-^ aspect has been analysed. The purpose of the present work is to review recent contributions to the problem of vroight estimation, to comment upon their applicability, and particularly to suggest siniple methods for use during the operations system investigation and initial project design stages,

It must be appreciated that the more simple a nx^thod for v^jight estimation Is the less accurate it is likely to prove, and hence, in general, a more elaborate method should be used to confirm initial predictions as soon as s\ifflcient Information becomes available,

2. Structure.

The aircraft structure constitutes a relatively large proportion of the t o t a l weight and i s t o a l a r g e e x t e n t under the c o n t r o l of the design teani. Although various methods are used for s t r u c t u r a l weight e s t i m a t i o n , probably the most usef-ul i s t h a t sometimes knavn as the '\7elght p e n a l t y ' concept. B r i e f l y t h i s c o n s i s t s of an estimate of the weight of a b a s i c o r ' i d e a l ' s t r u c t u r e , t o g e t h e r Yri.th the p e n a l t i e s

Incurred i n transfonnlng i t t o an a c t u a l s t r u c t u r e , i . e . allovirance f o r c u t o u t s , j o i n t s , c o n t r o l s , e t c . This method r e q u i r e s a f a i r l y d e t a i l e d laicwledge of the a i r c r a f t l a y o u t , ond con only be applied s a t i s f a c t o r i l y i n the l a t e r p r o j e c t design s t a g e s . Previous t o t h i s i t i s ncccrjsary to r e l y upon e m p i r i c a l , or seml-cni[jlrical formulae based upon p a s t exi^erience.

2.1.

In a conventional aircraft the lar^st individual contribution

to the structural weight comes from the wings. Because of this,

and the fact that the structural function can be defined fairly sinrply,

wing weight estimation has received much more attention than any other

component. The bibliography provides reference to nijmcrous works

concerned T/ith wing weight estimation rojiging from simple results

derived from statistical analysis to theoretical methods requiring

a detailed knowledge of design,

For a given all up weight, wing weight depends upon various

factors. The most important axe v/ing loading, thickness to chord ratio,

asisect ratio, taper ratio, svroeiDback, design diving sxxicd, maximum

normal acceleration ond relief loads. During investigation of the

operations system stage of an aircraft design, '.vhere the main object

is to decide ui^on the best type of aircraft to suit the system end

to formulate requirements, only a vague idea of these parameters is

llkelly to be available. Probably the most Important single parameter

is the wing loading and the following simple formula gives wing weight

in terms of wing area and all up weight.

r^t-^f ' y ^

= 0 ( 2 / + 0.08^7) . , . (1)

I

C is a f-unction of the ' design efficiency' , end has the following

gpproxlmate values.

C = 0.8 for a structure having no discontinuity or extreme

t/ or aspect i^atio, but having a large relief load,

G = 1 ,0 for an 'average' Vdng,

G = 1,15 for a wing structure \7ith nijmerous cut outs, lov7

t/ or very high aspect ratio, folding wings, etc.

c

It will be immediately appreciated that the accuracy of the

formula, which is based upon an analysis of a laxge number of existing

aircraft, is dependent upon the value of C used. As is alvro.ys true

in the art of weight es.tlmation, experience must jplay e. large part,

7

-It is Interesting to note, however, that an adverse value of one parameter usue.lly occxors with a correspondingly improved value of

another paraia^ter which, pojrtially at lee-st, tends to offset the first effect. For exanTple, the eissociation of low aspect ratio with very thin wings. Because of this, ve-rie.tions from tlie

'average' exc not generally unduly groeft,

The influence of the various paraijx;tors is such that the wing structure may be designed by either strength or stiffness

requirements, or in certe.in cases by both over different parts of the span. In this last case it appears that the structural weight is very similar to that necessary to meet the more dominant requirement over the whole span, A more accurate formula thoji tho.t given in Eq,(l) vshich allov/s for variation of the iniportent parameters when these become known in the early project stage is

:-¥^ = k[^^^^-^-(V°) ^ ciW. X 10^ + D -H 0 , ^ J . , , (2)

vdiore D = P,b,mir x l 5 ^ JAsec p ^ j ^ ^^^ / _^ ^^

_if

the wing is designed on strength

^^'^ I or D = 3L.b^ co£-v* x lö''^ j - ^ - ^ — j if

I . 2 J ... .;,10 [ J D

(Vc)( 1-0,1 6 6 M COS./-) the wing is designed on torsional stiffness.

(1)

This foraiula is based on one suggested by Grinsted^ ' in 1948 but he-s been modified to incorporate a later flutter stiffness

(19) •

requirement^ ' ond certain other refineuxjnts, F and L arc functions incorporated to make an approximate allov/ance for kinetic hee^ting effects, H is a function of aspect ratio and svvoepback, and g of inertia axis location.

The first term in Eq,(2) estimates rib weight and the last the wei£^t of leading ejid trailing edge fairings etc. The term incorporating the function Q is en allowance for discontinuities, joints, cut outs etc,

Q = 1.5 for on unbroken structxu:al box, incorporating bonding etc,

Q = 2,0 for an average design with no large cutouts,

Q = 2.5 - 3.0 for a design with large cutouts, folding v\dng etc, (The higher value if the fold is near to the root), On lo"v7 speed aircraft where the spexs react most of the

bending loads ond the skins only provide stiffness, it is necessary to add both terms ^for D Into the formula,

The other symbols are defined in the notation,

Applioation of the formula to a mrniber of recent aircraft wings has resulted in agreement of the order of - ^ being obtained,

but more evidence of accviracy is desirable, especially as it does not cover the aileron reversal co-se,

It should be noted that both Eqs. (l) and (2) assume that the wing bending moment is not transferred to the body, i,e. the

structural box is unbroken across the centre section. If this is not the case, the wing itself will be lighter but the fuselage will suffer a consequent penalty, the value of v/hlch is discussed

later, (see Eq,(5)), The total wing and fuselage weight is approximately Independent of the type of joint,

As soon e.s the v/ing desi(;'7i ond layout has been finalised, it is desirable to check the estimated wing weight by a more elaborate method. Such a method consists of estimating the weight of each structural coniponent required, together with the weight ixjnalty allovrance „ Examples of such methods are tjiose of Ripley (5), Burt (8), Micks (3), Hyatt (6) and Hamltt (7)^, .Of

_ 9

-a i r c r -a f t designed to meet B r i t i s h requirements,

2 , 2 , Fuselage,

The fuselage p r e s e n t s e. much more complex x^roblem than the vong

a s the weight de;ipends t o a f a r g r e a t e r e x t e n t upon d e t a i l

c o n s i d e r a t i o n s , p a r t i c u l a r l y cut o u t s , vdjadovTs end doors. In order

t o obtain a \TCight estimate m t h the desired accuracy i t i s

necessary to analyse a fuselage i n d e t a i l ,

I t woiold apxoear t h a t there i s a connection betv/een fuselage

weight and voluino, or surface a r e a , for a given a l l up weight, "v

although t h i s r e l a t i o n s h i p i s much l e s s defined than i n the ce.se of

wings. The following r e l a t i o n s h i p i s suggested as being s u i t a b l e

for an i n i t i a l estimate of fuselage weight,

Wp = a S p + /9W , . . (3)

v/here a = 0,4 and /? = 0,062 for a passenger a i r c r a f t ,

a = 0 , 9 and /S = 0.062 for a f r e i g h t e r or t r a n s p o r t a i r c r a f t ,

a = 0,65 and /? = 0.062 for f i g h t e r s , bombers and t r a i n e r s ,

a = 0 , 6 5 end /5 = 0.085 for f l y i n g b o a t s .

a = 0 , 4 caid /? = 0,038 for a i r c r a f t with nose p i s t o n engines

Again it vd.ll be seen that experience is desirable in the application of this formula, and at best only a rou^!;la approximation suitable for the oxxjrations system analysis is obtained,

A more accurate method of analysis, which requires a more detailed knowledge of the f-uselage as well as considerable experience is tho,t proposed by Ripley (12) , The weight of a 'basic fuselage' is calculated from the formula

:-Wp =

6,3

Vj^°*5 j 1,85 + ^ 1 Sp''*^ . 10^ lbs. , , . (4)

^ B+h -^

Driggs (14) makes an attempt to correlate nacelle weight with engine performance, but the evidence is very scenty,

Pod mountings usually weigh about 0* 18 of the engine weight, 2.5. _Tail Booms.

Althougli t a i l booms a r e s t r u c t i j r a l l y good, t h e y have a s m a l l d e p t h i n comparison t o t h e l o a d t h e y a r e r e q u i r e d t o c a r r y and a r e c o n s e q u e n t l y h e a v y , A f i g u r e of 0,025^/ t o 0,03W o r

5 - 6 l b / s q , f t i s t y p i c a l . The f u s e l a g e a s s o c i a t e d w i t h t h e t a i l booms w i l l be l i g h t e r t h a n t h e c o n v e n t i o n a l c o n f i g u r a t i o n and t h e

t o t a l w e i g h t of booms and f u s e l a g e vd.ll n o t v a r y g r e a t l y from a c o n v e n t i o n a l a r r a n g e m e n t , p r o v i d i n g t h e f u s e l a g e does n o t e x t e n d an a p p r e c i a b l e d i s t a n c e b e h i n d t h e v d n g ,

The performance of an u n d e r c a r r i a g e i s d e f i n e d by r e l a t i v e l y simple p a r a m e t e r s and hence i t i s amenable t o t h e o r e t i c a l t r e a t m e n t when knowledge of t h e geometry and d e t a i l s a r e a v a i l a b l e ,

U n f o r t u n a t e l y hovrover, t h i s i n f o r m a t i o n does n o t become a v a i l a b l e u n t i l t h e r e s t of t h e a i r c r a f t i s d e c i d e d upon ond f o r p u r p o s e s of

system e v a l u a t i o n i t i s n e c e s s a r y t o base e.ssumptlons upon s t a t i s t i c a l r e s i ü - t s . The most i i n p o r t a n t f a c t o r i n t h e d e s i g n of t h e u n d e r c a r r i a g e a p a r t from t h e a l l up w e i g h t i s t h e v e r t i c a l v e l o c i t y of d e s c e n t , T h i s i s p e x t i c u l a r l y h i g h f o r n a v a l a i r c r a f t and hence t h e s e u n d e r c a r r i e . g e s a r e r e l a t i v e l y h e a v i e r , ¥ j ^ = 0.037V/ ( 0 . 0 4 4 7 f o r n a v a l a i r c r a f t ) "^ W = O.OO7W (O.OIW f o r n a v a l a i r c r a f t ) \ , . . ( 8 ) Y7 "TU

= O.OO3W (0,005Y/ f o r naval a i r c r a f t ) j

These valvies are only apxsroximate and Ttdll ve-ry vd-th parameters such as undercarriage l e n g t h , r e a c t i o n f a c t o r , rake and t y r e

p r e s s u r e , ond i n t h i s connection experience must play a b i g p a r t .

An attempt by Biort and Ripley (I6) t o evaluate 'undercarriage weights more thoroughly has proved u n r e l i a b l e i n r e c e n t a p p l i c a t i o n s

13

-due to the f a c t t h a t i t was based on low tyre j)rossures end

outdated requirements. The sit\aation has been remedied by a

r e c e n t r e p o r t of P h i l l i p s ( 1 7 ) . I t vd.ll be appreciated the.t a t

the i n i t i a l p r o j e c t desif^n ste^ge, tyre and v/heel s i z e s are often

not a v a i l a b l e and thus the f i r s t p a r t of the weight estimation

process i s t o decide upon these d e t a i l s . P h i l l i p s p r e s e n t s some

curves giving tyre load carrying capacity i n terns of the product

of tyxe diameter ond width for an 'average' range of t y r e s

(w = 0,29d), and various tyre xoressures. Tyre plus tube vrolght

follows as a function of the same parameters. The v/heel size i s

estimated from the average value u = 0.48d ond the v/eight of the

wheel i s given i n terms of t h i s . The brake weight i s given as a

fvinotion of the k i n e t i c energy per b r a k e , and the t o t a l wheel

u n i t vre'ight i s the sum of the three q u a n t i t i e s . Normal and

emergency cases are considered,

As an a l t e r n a t i v e t o t h i s procedure F i g s , 2 - 6 can be used,

F i g . 2 gives the required tyre diameter for a given load and tyre

pressure ond i s based on a p a r t i c u l a r family of tyres which meJkes

allowance for the tendency of high pressure tyres to have reduced

w i d t h s . The values of these vddths and the rim diameters are given

i n F i g s , 3 cind 4 . Should i t be necessary'- to use a vdieel of reduced

width the new diameter should be estimated on the b a s i s of a constant

product of width and diameter but the subsequent operations should

be based on the o r i g i n a l diameter. The s t a t i c load must not exceed

one t h i r d of the maximum dynamic load U,

The v/elght of the tyre plus tube i s estiiaated from F i g , 5 ^^d

the brake vreight from F i g , 6, This l a s t figure i s given as a

function of a i r c r a f t s t a l l i n g speed and a drag f a c t o r f.

f = 1 ,0 for a very clean nosevAieel a i r c r a f t vdth

no drag f l a p s ,

f = 0,7 f o r a tailwheel a i r c r a f t vdth drag f l a p ^ vdth

intermediate values i n between.

14

-The t o t a l wheel i m i t vraight i s then given by

For braked v/heels : - 1 ,43 (W,p + y^) per v/heel j

For tmbralced vdieels : - 1 .6 Y/™ per virheel /

(9)

stmcture in terms of the resioltant factored load normal to the undercarriage leg, P, Thus this value takes account of reaction factor and rake. The curve given is approximately defined by

:-0.00127X, P°-''^ . . . (10)

where 6 i s the distance p a r a l l e l to the leg from the ground

1 Ine vdth the l e g extended to the point where the xjndercarriage

load i s t r a n s f e r r e d t o the airframe ( i n I n c h e s ) .

The t o t a l weight of the undercarriage i s the s\mi of the v*ieel

and s t r u c t u r e vreights plus an allovrance of 2^ for miscelloneous

paxts and a f u r t h e r 5^ i f the u n i t i s r e t r a c t a b l e . The estimate

does not include the weight of bogies v;hich amount t o axTproximately

O.OO3W, ond vAien these are used t h i s must a l s o be added i n to get

the t o t a l undercarriage vreight.

2 . 7 , F l o a t s ,

There i s very l i t t l e evidence available on f l o a t vraight, but

what . there i s i n d i c a t e s the follovdng values for f l y i n g boats :

-For fixed f l o a t s : - O.OIW ^

For r e t r a c t a b l e f l o a t s : - 0.015W J . . . (11)

2 . 8 . F l a 2 s ,

This i s Included i n the vdng v/eight, x^aragraph 2,1 but f o r

i ' d e t a i l e d estimates see Liort ( l 8 ) ,

3 . 0 . Povrer P l a n t .

I t i s normal f o r p r o j e c t designs to be based on engines which,

a t l e a s t , are f a i r l y advanced paper studies and thus a power p l a n t

weight i s a v a i l a b l e , together vdth c e r t a i n a c c e s s o r i e s . During the

operations system a n a l y s i s however, i t i s convenient t o use more

15

-general Information and therefore typical values of

^ovier

plant

weight are given as fijnctions of performance for recently developed

engines.

I W.

Turbojet :- 0,2T

Prcpjet :- 0,5(E.PIJ'.)lbs (excludes propellers) /-.„N

Piston Engine :- 200 + 1.04 (H.P.)lbs (excludes

propellers)

These figures are based on maximijm performance, but exclude

I special boosting, e.g. methanol injection and reheat,

Propeller weight is normally approximately given by

:-/ 0.24 ( H , P . ) lbs. . . . (13)

f^~~\ \ The engine i n s t a l l a t i o n vrelght i s usually of the order of

C 10^ - 2C^ of the povrer p l a n t vrelght. .. •

4 . 0 . Systems.

The aircraft systems, l,e, flying controls, hydraulics,

electrics, de-icing, pressiorisation and air conditioning, and

fuel system vary considerably from one aircraft to another, and

it is a very difficult problem to attempt an accijrate v/olght

estimate. The best method is to use the values of an existing

similar type, but unfort-unately these are not alv/ays available.

Certain very approximate suggestions ceji be made to coArer this case.

The flying control system is taken to Include flap operation.

The vreight of this system is given approxime.toly by

For large airliners end transport aircraft

:-35 + 0.008W

(14)

For other types : - 35 + 0.00517

These f i g u r e s do not include powered oxxsration of the

c o n t r o l s . The l i t t l e evidence a v a i l a b l e i n d i c a t e s t h a t when

t h i s i s used the term Independent of \i should be increased t o about

100. Large trejisport a i r c r a f t are usually equipped vdth complicated

f l a p s and t h i s accounts for the liigher weight,

4 . 2 . Hydraulics, Pneimuxtics end E l e c t r i c s .

The v/elght of a hydraulic system n a t u r a l l y depends upon the

number of services operated. The sams ax^plics to pneumatic and

e l e c t r i c a l systems. To some extent they are i n t o r - r e l a t e d i n

t h a t f o r a given type of a i r c r a f t the t o t a l number of services

operated v d l l be s i m i l a r . For t h i s reason they are considered

t o g e t h e r h e r e , but even so there i s a large variixtion i n apparently

s i m i l a r t y p e s . Exclusive of d e - i c i n g and assuming t h a t every

e f f o r t i s made to reduce weight - e , g , high pressiire hydraulic

system, llghtvraight cables e t c . , the t o t a l weight of these povirer

systems usually f a l l s i n the range "Zfo - h^o of the t o t a l a i r c r a f t

weight. Any sxx3cls.l feature's, such as a complex r a d a r i n s t a l l a t i o n ,

v d l l of course tend to increase t h i s v a l u e .

4 . 3 . De-icing,

Again there can be considerable v a r i a t i o n , b u t f o r f u l l

d e - i c i n g of the airframe, the penalty i s about 0,8$'S of the t o t a l

vrelght or :

-500 + 0,003Y/ . . . (15)

which ever i s the lower.

• 4 . 4 . Pre,ssurisation and Air Conditioriing,

I n s u f f i c i e n t evidence i s a v a i l a b l e t o enable any recommendations

t o be made for high speed a i r c r a f t which experience k i n e t i c h e a t i n g ,

Each case must be t r e a t e d on i t s m e r i t s ,

For passenger aircra^ft the vreight of the equipment i s

approximately 29? of the a i r c r a f t vrelght or :

-700 + 0,002W + l o o t , , , (16)

which ever i s the lower,

17

-(18) 4 . 5 . F u e l System.

The v/elght of the f u e l system i s v e r y dex^endent upon t h e number of power p l a n t s and t a n k s and t h e i r r e l a t i v e l o c a t i o n s . Hence i t i s v e r y d i f f i c u l t t o g e n e r a l i s e . The f o l l o v d n g f i g u r e s I n d i c a t e t h e o r d e r of f u e l system w e i g h t t o b e e x p e c t e d :

F o r c i v i l a i r l i n e r s and t r a n s j j o r t a i r c r e d ' t :

-0,0l6W - 0.02W ") _, ' ^ /^j\ F o r o t h e r t y p e s up t o : - O.OJW j

depending iipon t h e coinplexlty of t h e s y s t e m ,

These f i g u r e s I n c l u d e t a n k w e i g h t ,

The vreight of t h e t a n k s t h e m s e l v e s caii be e s t i m a t e d from :

-For p l a i n m e t a l t a n k s 10 + 0.69V ^^ Crash p r o o f t a n k s ( m i l i t a r y f l e x i b l e ) 6,3 + 0.38VV F l e x i b l e c i v i l t a n k s 10 + 0.12V * * Drop t a n k s 1.ÖV tdiere V i s t a n k c a p a c i t y i n g a l l o n s . The t a n k p l a t i n g f o r f l e x i b l e t a n k s u s u a l l y weighs a b o u t t h e same a s t h e t a n k .

R e s i d u a l f u e l and o i l average a t 3Sfo of tanlc w e i g h t ,

5 . 0 , Equipment ejid m i s c e l l a n e o u s i t e m s ,

These iteiiis can amount t o a s u b s t a n t i a l p r o p o r t i o n a l of t h e t o t a l w e i g h t and axe o f t e n v e r y d i f f i c u l t t o p r e d i c t . I n t h e c a s e of m i l i t a r y a i r c r a f t i t i s e s s e n t i a l t o use e x p e r i e n c e g a i n e d from a comparable t y p e o r p r e f e r a b l y tlie a c t u a l r e q u i r e m e n t s v/hen

t h e s e a r e a v a i l a b l e . I t i s p o s s i b l e t o be a l i t t l e more d e f i n i t e i n t h e case of c i v i l a i r c r a f t .

5.1 . S e a t s and F u r n i s h i n g s .

E j e c t o r s e a t s of the o l d e r type weigh a b o u t 200 l b s and t h e more r e c e n t l i g h t w e i g h t ones a b o u t 100 l b s . The o r d i n a r y crew s e a t s have a w e i g h t of ax^proximately 30 l b s and t h e s p e c i a l l i g h t

-welfiht military transport type 18 l b s .

The v/eiglit of yjassenger seats in civil aircraft varies according "to the tyi:>e and duration of flight. On aircraft intended for short fliglits, the seat weig^ht i s usually about 25 ibis but for longer fliglits 35 lbs i s nearer the average value. Specie.l lightweight scats of recent design Vv-eigh as l i t t l e as 21 lbs per x-?assenger for reasonable luxury.

I t is often convenient however to consider seat and

fiornishing weight together. A n analysis of recent large airliners ^ g g e s t s the following values for total fiomishing wisight :

-First class accommodation : - 60 + 12t lb/passenger^

Toiorist accommodation : - 25 + 12t Ib/passengerj • • • ^ J

The higher f i r s t cle.ss figure is not only a function of greater luxury, but also of lower seating density. The exclusion of seat weight axTpears to have l i t t l e effect xipon the time variable term and yi.elds : - ("^

First class accommodation : - ,'3<?+ 12t lb/passenger | (on") Toiorist accommodation : - 35 + 12t lb/passenger\

5.2. Instruments.

The vroight of instruments on a small trainer or communications aircraft is of the order of 40 lbs rising to about 60 lbs on the more advanced types,

In the case of airliners the v/eight is usually in the region of 350 lbs - 450 lbs mainly dependent upon reoige,

5.3. Radi o end Radar.

A siraxjle rexlio installation inciors a direct v/eight penalty of eJbout 35 lbs o On more advanced aircraft of the trainer type the radio installations vdll probably weigh of the order of 200 lbs. The Installation in airliners is more extensive ejid accoijnts for 800 lbs - 1000 lbs, the higher value a-pplylng if some

- 19 - .

form of radar is used.

5,4. Fire Precautions.

The fire precaution v/elght naturally depends very much on the engine end fuel system layout, but an e.verage value for civil airlines is

:-(0.1 + 0,1s)W X 10""^ . . . (21) where s is the himiber of engines.

The follovdng figures are based on recent lightweight oxygen containers

:-Y/t. for one crew member for four hours 55 lbs, )

, , , (22) Yrt, for one passenger for four hours 20 lbs. i ' * *

5.6. Pednt.

P a i n t v/eight can vary considerably according t o typo, number of c o a t s , e t c . , but an average value i s 0.035 I b / s q . f t . of s u r f a c e ,

R e f e r e n c e s ^

1 . Simple Formiilae f o r X3redictin.g the vreights o.f Wing, F u s e l a g e and T a i l U n i t S t r u c t u r e s ,

F , G r i n s t e d o

R,A,E. R e p o r t S t r u c t u r e s 24 (May 1 9 4 8 ) ,

2. Aircraft Wing Weight Estimation,

J, F, Caxreyette,

Aircraft Engineering Jan, 1950, 3. Stxactural Weight /oaalysis - Wing Equations.

W. R. Mrlcks,

P r o j e c t Rand Rep.R198 ( D e c . 1 9 5 0 ) . 4 . The E s t i m a t i o n of Wiiig W e i g h t .

J . S o l v e y ,

Aircraft Engineering May 1951. 5. A method of Wing Weight Estimation.

E. L, Ripley.

R.AoE. R e p o r t S t r u c t u r e s 109 (May 1 9 5 1 ) .

6 . A method of E s t i m a t i n g Y/ing Yfeights, A. H y a t t .

31oumal of A e r o n a u t i c a l S c i e n c e s

June 1954.

7. structural Weight Estimation by the Weight Penalty Concept. R . L , HaiiTtrdtt.

S .A oW,E, P a p e r 141 ( 1 9 5 5 ) . 8. Weight Prediction for Wings of Box Construction.

M. E. Burt.

R.AJü, R e p o r t S t r u c t i x r e s 186 * (Aug. 1 9 5 5 ) .

9 . The Yföight of F u s e l a g e s f o r C i v i l A i r c r a f t . L . Y*'. R o s e n t h a l ,

S.3.A.C. Weight Control Panel Rep. (July 1948).

1 0 . S t r u c t u r a l Y/eight / j i a l y s i s - F u s e l a g e and S h e l l S t r u c t u r e s .

'I W. R. M i c k s .

: P r o j e c t Rand Rep.R172 ( J a n . 1 9 5 0 ) . 11 . The Y/eight Asx^ect i n A i r c r a f t D e s i g n ,

L . W. R o s e n t h a l ,

Journal Royal Aeronautical Society Mar, 1950,

1 2 , A Method of F u s e l a g e Structxnre Weight E s t i m a t i o n , E . L . R i p l e y .

21

-P r e d i c t i o n of Fuselage and Hull Structure Y/eight.

M. E. Burt and

J , P h i l l i i i s .

R..A.E. Re-Dort Structures 122

(Ax^ril 1952).

14. A i r c r a f t Design Analysis.

I . H, Driggs.

' Joiomal Royal Aeronautical Society

Feb, 1950 o

1 5 . A Sample Method of T a i l Unit Structure Y/eight Estimation,

E, L, Ripley,

R.AJl, Report S t r u c t u r e s 94

(Nov.1950),

16. P r e d i c t i o n of Undercarriage Y/eights.

M. E, Burt and

E, L, Ripley,

R,A^E, Report Structiores 80

(June 1950).

17. A Method of Undercarriage Y'eight Estimation.

J , P h i l l i p s .

R.A.E. Report Structtires 198

(Mar,1956),

1 8 . P r e d i c t i o n of Aileron and Flap Y/eights.

M, E . B u r t .

R.A.E, Report Structures I I 6

(1951).

19. Aeroelastic Problems i n Connection vdth Higli Speed F l i g h t ,

E, G, Broadbent.

JouiTial Royal Aeronautical Society.

J u l y 1956,

Appendix 1 .

S t a n d a r d Y/eights of Crew and P a s s e n g e r s .

M i l i t a - r y 200 l b s iper meji ( i n c l u d i n g x^exachute b u t e x c l u d i n g p r e s s u r e s u i t , e t c . ) . C i v i l J3aggage f o r r o u t e s I n t e m a i t È « » ^ C o n t i n e n t a l Overseas P a s s e n g e r : Male I65 l b s p l u s 33 lb)s o r kk l b s o r 66 l b s FemeJ.e 143 lb>s p l u s 33 l b s o r 2f4 Ibis o r 66 l b s C h i l d 85 l b s p l u s 33 l b s o r 44 l b s o r 66 l b s ( 2 - 1 2 y e a r s ) C h i l d 17 l b s p l u s 33 l b s o r 44 l b s o r 66 l b s (Under 2 y e a r s )

Crew: Male I65 l b s p l u s 22 l b s o r 33 l"bs o r 44 lb>s Female I 4 3 l b s p l u s 22 l b s o r 33 lb)s o r 2(4 lb)s

/

FKTOR k

0-4 0 6 X TAPER RATIO

FIG. I. TAPER CORRECTION FACTOR.

3 U y <

i

J

1

}

1

7/

'h

7

TYRE WIDTH U INCHES 14 12 lO e 6 4 2 y

A

y

p

i

y

TYRE DIA. d INCHES

FIG. 3. TYRE WIDTH

WHEEL RIM OIA U. INCHES

FIG A WHEEL RIM DIAMETER

WT

lbs.

o 25 3 0 35

TYRE DIA. d INCHES

FIG. 5. TYRE WEIGHT Wy

. ^ ^

0^

A

r

A

V

//A

V

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