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ENGINEERING REPORT NO.

FC DERIVATION OF WEIGHT Rp 3ERMS OF

PARAMEORIC DESICai AKALYSIS TOR

PROPELLOFLANE TRaNSPORT 3TUDY

Contract !Jonr 1657(00)

ALGl

m 9 94 /-vsy

NOTICE; THIS DOCUMENT CONTAINS INFORMATION AFFECTING THE

NATIONAL DEFENSE'OF THE UNITED STATES WITHIN THE MEANING

OF THE ESPIONAGE LAWS, TITLE 18, U.S.C., SECTIONS 793 and 794.

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ANY MANNER TO AN UNAUTHORIZED PERSON IS PROHIBITED BY LAW.

y

H« Feldman SZk£6 MILLER HELICOPTERS CMCOKCO TiTte: Derivation of Vfeight Hp Terms of

Parametric Design Analysis For Propelloplane Transport Study

Contract Nonr 1657 (00)

1016

■ CFOHT NO

CONFIDENTIAL

INDEX

Page No.

LIST OF SYMBOLS ii

SUMMARY 1

INTRODUCTION 2

DERIVATION OF » TERK 5

A. Pov.-cr Package • • • . 5 5. Propellers P C. Wintrs 12 D. Fuselage 13 E. Landing Gear 13 F. Empennage ................ 13 G. Supplemental Control System lU H. Wing Tiltinr Mechanism 15 I. Fixed and Operational Equipment 17

REFERENCES 21

CONFIDENTIAL MH)» e

•■«»•■»o H. Feldnan | $/h/S6 HILLER HELICOPTERS 11

CMCCKCO T.TLi- Derivation of tfeight RF Terns Paranetric Desim Analysis For Prcpelloplane Transport Study

• oon. IOJ45

mtwamr NO l«.

CONFIDENTIAL Contract Nonr 16^7 (00)

LIST STtMBCfLS

Ap - Propeller Disc Area, ft2

AF - Propeller Blade Activity Factor

Aq - IxP Propeller Excitation Factor

AR - Aspect Ratio

B - Number of Propeller Blades

b - Wing Span, Feet

c - Wing Chord, Feet

Dp - Propeller Diameter, Feet

HP - Installed Normal Rated Horsepower, StandaM Day, Sea Level

Kp - Ratio of HP/Thrust

MQ - IxP Propeller Blade Root Bending Moment, Ft.-Lb.

R - Ratio of Weirht to Aircraft Design Gross .eight

Q - Torque, Ft-Lb.

S - Wing or Tail Surface Area, ft2

Vf - Propeller Tip Speed Ft./Sec.

W - Weight

wjj - Hover Disc Loadinr:, Defined as Hover Thrust/Ap, Lb./ft2

W/S - Wing Loading, Design Weight/Wing Area, Lb./ft2

CONFIDENTIAL 60-030 e

NAME

H. Feldman g/li/gb MILLER HELICOPTERS T 1

CHCCKCO

APPROVCD

Derivation of Weight Rp TeiT'S c Parametric Desi-m Analysis For Prcpelloplane Transport bt';o:'

MOOKC ioh?

Contract Koiir Ibf? (O""') CONFIDENTIAL

SUMMARI

This report sujmnarizes the structural design criteria and presents the derivation of the weight Rp equation for parametric determination of the desiftn parameters of the minimum RTOES veipht aircraft capable of fulfilling the performance specifications of Contract Nonr 165? (00).

CONFIDENTIAL MM)» e

H. Feldmar: . g/jjA6 MILLER HELICOPTERS CHCOKKO Derivation of Weight Rp Terms of

Parametric Design Analysis for Prooelloplane Transport Study

10U8

»•li7U.5

CONFIDENTIAL Contract Nonr 1657 (00)

INTRODUCTION

The Weight Rp Equation

Use of the Hiller Rp method of parametric optimization for the specified transport propelloplane mission requires the develop- ment of an analytical expression for the variation of the ratio of fuel weight to gross weight which is permissable at any gross weight in terms of the var_ .bles of the investigation.

In general

where

Wp + Wc + WE + Wp ♦ WpT

wc

Wc

Gross Weight

Crew Weight

Wp • Design Payload

WE - Empty Weight Less Fuel Tanks

Wp - Allowable Fuel V^j - Fuel System Weight Weight

re-writing Wp + WpT - Wg - Wp - wc - WE

and placing into ratio form by dividing by gross weight

Rp ♦ Rpx ■ 1 - Rp - RQ - ■

where fi is designated as the ratio of empty weight,less fuel system weight,to gross weight. To ensure compatibility between fuel and fuel tank weight, tank weight is assumed to be proportional to the amount of fuel stored, hence to fuel weight.

Rp (1 + kp) - 1 - Rp - Re - »

Therefore, the Rp can be expressed

RF - (W kp) U " »F - 8C " ■)

This equation is the generalized weight equation of the aircraft and is referred to as the Weight Rp Equation. The Rp parameter provides

the common link between weight and aerodynamic characteristics.

CONFIDENTIAL 60-OX c

H, Feldman $M5k HILLER HELICOPTERS CHCCKCO

APPROVED

nTLt: Derivation of Weight lip Ttms of Parametric Design Analysis For Propcllonlane Transport ^»tudy

10li8

"TWO hlU.S

CONFIDENTIAL Contract Nonr 1657 (Oü)

For the mission of this contract, the design payload is specified as 'n000 lbs,, a cre-w of three roighs 60G lbs., and the weight of self sealing tanks is assumed to be ,9 lbs, per rallon of fuel.

Thus Rp - "000 A^g

Rr; « 600/.'g

1 + Iqr = (1 + -i2_) 1.15

hence Rp 1.1^

(1-1^22-1)

The remaining unknown, I, is the sun of the weight ratio expressions of all individual components which comprise the empty weight of the aircraft and is, therefore, a function of the desim parameters af- fecting each.

Five parameters are chosen as the fundamental variables of the study. These are:

1. Disc Loading

2. Tip Speed

3. Aspect Ratio

U. Wing Loading

Gross Weight

The na,ior effort of the weight analysis is, therefore, the derivation of the weight expressions for the conFonent items of 5 in terms of these five variables.

Structural Criteria and Weight Prediction Approach

By nature of the broad scope of the parametric analysis utilized in this study, establishment of structural design criteria is limited to .Ter.erali- zations sufficient to insure realistic weight estimations of the aircraft components whose weights are a function, in some manner, of the aircraft leads.

Design loading for propeller blades is established to be trie more critical of the IxP vibratory moments occurring during transition or normal fixed wing flight regimes.

CONFIDENTIAL 60-010 e

MAMI 1 H, Feldnan

CMCCKCO

AP^aovco

5A/$6 HILLER HELICOPTERS Derivation of i/ei^ht Rp Terms of Parametric Design Analysis for Propelloplane Transport Study

IOI48 RE PORT NO.

CONFIDENTIAL Contract Konr 165? (OC)

Airfrarric arui v.'ings ars designed to +5«0, "3<.';;g altinate load fac.-r.rs in order to provide a general strength level aaequately representative of syruTietrical maneuver and landinr conditions for aircraft- of this siae and function.

Design loading for the viing tilting mechanism, which is most critical during an asymmetrically braked forward landing roll with wing tilted to vertical position, is approximated by an equivalent 2,63g lead fac- tor applied forward through i.g. of hinged mass when the wing is vertically positioned.

The approach to the probler. of practical weight prediction considers the fact that, in general, design requirements for most of the air- craft components are similar to those of current conventional air- craft, and the weights of these components can be most practically expressed by e.rpiricai equations derived from data on operational aircraft with similar design parameters. These weights are quite representative of current design practice.VJherever required for com- ponents which are peculiar to this type of aircraft,or represent unique applications of conventional components,, detailed treatment on a more analytical oasis is accorded.

CONFIDENTIAL 40-ox e

NAME

H« Feldman CHCCKCD

APPROVED

5A/56 HILLER HELICOPTERS r.TLE: Derivation of Weight Hp Terms of

Parametric Design Analysis for j 10U8

Propelloplane Transport Study | "«^Q»™" UTh'S

CONFIDENTIAL Contract Nonr 1657 (00)

DERIVATION OF THE 5 TERM

The components comprising the empty weight items of this aircraft are divided into the following nine groups.

A* Power Plant (Including engines, transmissions, engine controls, accessories, engine mounts, vibration dampers, nacelle.)

B. Propellers (Excluding propeller controls, anti-icing and spinners.)

C. Wings

D. Fuselage

£. Landing Gear

F. Empennage

G, Supplemental control system (includes auxiliary engines, ducts, jet deflecters, fuel, etc.)

H. Wing tilting mechanism

I. Fixed and Operational Equipment (Includes surface control systems, hydraulic, electrical, pneumatic systems, furnishings, navigation equipment, anti-icing, and air conditioning provisions, electron- ics, etc«)

Expressions for each group are derived individually below, and the sum of the expressions defines 31.

A. Power Plant Weight Ratio

Power Plant Weight includes the weight of the engines, transmissions, engine controls and accessories, engine and transmission oil, and oil systems, engine mounts, vibration dampers, firewalls, and nacelle cowling.

The gas turbine enenn«s for this aircraft are similar in construction to the current Allison T-ItO engine, in that there are two power sec- tions geared to a common transmission in each nacelle. From the gen- eralized specific weight curves for engines forecast for 1965 (Reference 1), the following relationship between engine weight and rormal rated power at sea level is derived for the ranpe of horse- powers indicated.

CONFIDENTIAL 6O-010 e

H. Felctaan ! 5/u/: HILLER HELICOPTERS CHCCKCO r.Ttc: Derivation of Weight ftp Terms of

Parainetric Design Analysis for Propelloplane Trajisport Study

(Contract Nonr 1657 (00)

IOI18 RCFOOTNO- I.T), r*

CONFIDENTIAL

W

^V nacelle

,-.26 = 2-0^ H^celle^OOO^HP^n^l^OOO)

hence, the veight ratio of the engine in one nacelle is

We . 2.02 (,HP-'2SP = 2.C2 ^j^

Total normal rated design t'nrust at sea level, as required for hover ceilin,: requirements, ■ T ■ 1.3 W~, so the HP of one nacelle of a four nacelle aircraft may he expressed

HP nacelxe Z IiPtot 1 u

HP T

1.3 MiP fj« 1.3 Xp w

vhere Kp ■ HP/T and HP ■ total normal rated horespovrer at sea level installed in the aircraft.

Hence, the engine weight ratio raay finally be written

We E

2.02 1.3 Kp W_ i..

.7) Ik _ fiQ V'lU •■-,2.6 - .ÖÖ Kp Wg

This expression includes v:eight of accessories and engine controls, but not reduction '-earing.

Transrdssions for this aircraft couple the two power sections cf each nacelle to the propeller shaft and are of the planetary train type with two inruts and a coaxial output. Statistical data for transmissions of this typ.?, which includes the weights of gearboxes and cer.trifural and overrunning clutchea, indicate the following relationship between transmission weight and maximum torque on the low speed output (References 2 and 3).

V'p .081 kn Q .88

kj^ a fro'oor to account for the number of inputs and outputs is evaluated TO be l.yO from studies of current transmissions of tnis type (T- L, T-56, T-uO).

CONFIDENTIAL 6O-0» e

H, Feldman 5A/g6 MILLER HELICOPTERS 1 "dkat

CHCCKCO

APPROVED

TITLI: Derivation of Weight i^Terns of Parametric Design Analysis for Prcpelloplane Transport Study RL^onr MO- hlh

CONFIDENTIAL (Contract Nonr 165? (00)

For one nacellej total design torque is assajned to be 75 percen-; maximuin torque available at sea level military pover. This de- rating of the transmission effects a significant weight saving in view of the larrre excess of oover necessarily available at sea level in order to meet the hover requirements of 6000 feet on a 95° F day.

thus Q - .75 550 HP Dp

2VT

Using the previous notation for horsepower/nacelle, military horsepower may be expressed

HP^^L - 1.12 H^A = 1.12 (±f2 Kp Wg)

and defining the hover disc loading as T/total disc area

Hence, the weight ratio of one transmission is

w: ■ 3'^^] w .32

WH .uU

Oil consumption for engines and transmissions is conservatively assumed to be 1.5 gallons per hour per nacelle, based on average requirements of present day engines. Oil tanks weigh approxi- mately two pounds per r-allon of oil, and turbine grade oil weight is assumed to be 7.9 lb/pallon.

Hence, the weight ratio of oil and oil system per nacelle may be expressed, for an assumed three hour mission with 100 percent re- serve

W oil + oil sys. (1.5)6(7.9 ♦ 2) 70 * 20 w "g

90 wl

Starter weight is neglected since the engines are started by the auxiliary power unit bleed air.

CONFIDENTIAL 60-on e

NAMC

H« Feldnan 5/U/$6 MILLER HELICOPTERS CHCOKCO

APPKOVCO

Derivation of Weight Up Terms of Parametric Design Analysis for Propelloplane Transport Study

lOhB

RCPORT NO. l|7Uo5

CONFIDENTIAL Jontract Nonr 1657 (00)

The nacelle weight includes engine mounts, oil cooling systcr". firewalls, vibration isolation systems, coulinss, etc. Present day installation weight averages about 68 percent of engine weight« Assuming that the twinned engine system increases in- stallation weight required f'-r a single engine by 50 percent, then

w; i^m-^w: or

WN

w g

.U5 Kp7U -.26

Total povrer plant weight ratio is the sum of the above expres- sions (one nacelle only).

S2 - 1.33 K^ü „--a* . |2 . }.h'K

wg " wß

W, .32

ua Kp, the ratio of total installed normal rated horsepower at sea level to design thrust is a characteristic of the propellers chosen and will "cc related to tip speed and disc loading in section B.

The above expression, of course, represents "rubber engines". ?or those portions of the study where hardware engines are required, weights are taken from the appropriate engine specifications.

3. Propeller Weight Ratio

Due to the severe vibratory loadings to which propellers for VTO aircraft are sub.-üct, propeller blade weight is best described in terms of the IxP loadings as follows: (Reference h)

WP - .0536 JM + .OOOeUU AF Dp2*3

q»* 'u

MQ is the critical IxP moment at the blade zero station. Thick- ness to width r.itio at the ail radius is approximately .lUP.

CONFIDENTIAL 60-0» c

NAME T ru/5b MILLER HELICOPTERS

CHCCKCO

APPROVCO

TiTi.1 Derivation of VJel^it H Terms of j MCJOIL 10U8 Parametric Design Analysis for | Propelloplane Transport Study «POIITNO [17^,5

(Contract Nonr 16^7 (00) CONFIDENTIAL

Weight of the nropeller hub for Troth single and dual rotation pro- pellers is apcroxiraately ?0 percent of total propeller weipht for the rapid pj tch change rates required vrith turbine engines. Hence, the weight ratio of the entire propeller may be written

Wp 1

B g

2B

r Mn 2 ^ j .0536 —70 ♦ .OOOeiili AFDp 2B

1 I .3

Defining hover disc loading as in the previous section, propeller diameter may he expressed

then Dp

\ U Tf WH /

and total weight ratio of one propeller is

/WAT) w;50. / MQ V-5 V.Tp

w ̂ = .1155 g ■)

This expression is equally valid for single and dual rotation Dro- pellers, since, in practice, D and AF are the same for forward and aft blades, the relationship between hub and blade weight is sim- ilar for both types, the B term accounts for the actual number of blades, and MQ ma-'r te assunied equal for forward and aft blades. Although calculations would show MQ smaller for aft blades due to straightening of the inflow, interference buffeting removes much of the conservatism of this assumption.

For s iraplification of calculations, the parameter MQ is used in place of MQ

B\vrH /

Hence

wg ^WH; L ^ wH J

CONFIDENTIAL 60-0'» e

PRCPAnCO H, Felcbnan \S/h/S6 CHCCKCO

HILLER HELICOPTERS

ARPPOVEC

Derivation of Weight Rp Terms of Para:netric Design Analysis for Propslloplane Transport Study

101:8

»CFOPT wo. u7 u5

CONFIDENTIAL Contract Konr 1657 (00)

In this expression MQ and B(AF) are not fundamental parameters of the Rp equation, but MQ*in the critical conditions can be related to disc loading and tip speed, and B(AF) will be optimized by special investigation to provide a unique value for each combina- tion of wy and V^,

Critical IxP Moments

Critical vibratory loadings occur in the follov:in^ flieht conditions.

1. Transition from vertical to horizontal flight.

2m Normal Airplane flight conditions.

Moments occurring during transition flight were investigated for all angles of the thrust axis fron horizontal to •"•ertical, using the data of Reference 5 for Mp, the pitching component, and the data of References 6 and 7 for My, the yawing component. It was found that good approximation of the critical moments can be ob- tained by simultaneous consideration of a yawing component arising from a yawing moment coefficient Cy = .0315» and the pitching moment occurring at a thrust line angle of 75° approximated by the product of propeller thrust and an arm of »193xpropeller radius

Moments arising from nom&l airplane flight conditions were cal- culated for critical values of "Aq" per methods of References 6 and 9.

Figure 1 summarises the most critical moments arising from con- sideration of tne two flirht conditions for the disc loadings and tip speeds investi prated.

B(AF)

Both power plant weight and propeller weight are functions of the propeller blade activity factor and the number of such blades, or B(AF): the former, because of the effect of E(AF) upon HP/'T and, therefore, upon power required; and the latter because of its ex- plicit appearance in the propeller weight equation.

Selections of a "best" value of this parameter through a separate optimization procedure is possible due to the fact that for this type of mission, within the range of tip speeds and disc loadings investigated, and the practical variations of B(AF) permissible, it is found that the effect of B(AF) upon the gross weight of the air- craft is due primarily to its influence upon the empty vieight of the aircraft rather tr.an its direct effect upon fuel consumption.

CONFIDENTIAL 60-030 e

VKCPAIICO I NAME

H. Felcbnai ihI5k \ HILLER HELICOPTERS CHCCKCO

«»paavco

T'T,"C: Derivation of V/ei^ht Hp Terns of Parametric DssifTn Analysis ,'or Prspelloplane Transport ->t'idr

11

Hf PORT NO )|7''.^

CONFIDENTIAL Contract Hour 1657 (0'))

Figure 2 sho-.-3 the combined weirjita of nov:erplant and propeller plotted versus B(AF) for single and dual rotation propellers for several disc loadings nt a representative; tip spervt and ^ross weight« Cutoffs rnpresentinr^ minimum pemissible values of B(AF) due to possibility of exceeding the allowable; blade utrt^sses (Reference U) or maximum permissible 1 lade stall are indicated. The B(AF) at which the combinsd '.reiFtht is a minimum in considered the optimum« Comparison of the mlnlmums for sinrle and dunl rotation propellers Indicate the host choice.

Variation of iiP/T v;ith B(AF) was obtained fron the propeller performance charts of Heference 10, and application of this opti- nisation procedure over the complete ran^e of disc lor rings and tip speeds yielded the following selections of E(AF).

Tip Speed Disc Loading B(AF) No« Blades I

000 ho U5o 3 60 710 6 PO P90 6 i

100 10-^0 O f

^00 iiO a95 3 ; 60 760 6 } ■o0 780 6 !

100 950 6 1 1000 iiO ■iao 3

60 530 3 "0 .Q?o ü ;

100 c,30 ö

Hence, v/eignt of the propellers Cor this aircraft is (ietermined as functions of the variables of the study«

Weights of spinners, prop anti-icers, and controls are included with fixed and operational equipment in Section I.

CONFIDENTIAL 6O-0TO e

H. Feldmar. DATE

CHCOKCO

t AFPROVCO

HILLER HELICOPTERS Derivation of Weight % Terrns of Parametric Design Analysis For Prcoelloplane Transport Study

12

10U8 »CFORT ND. I4.fl^m3

CONFIDENTIAL Contract Nonr 16^7 (00)

i C. Wj lug VJeif;h Ra

The wing chosen for this study is of conventional aluirdnum alloy sheet and stringer construction with spars located at the 15/6 and 50% chord. Planform and thickneas taper ratios of 211 are assumed; the wing is equipped with leading edge slots and trail- ing edge simple type flaps, and is hinged at the rear spar for tilting. Although the aft $0% of the wing is not continuous across the fuselage, the structural box is not interrupted. For weight purposes a symmetrical 15^ airfoil is assumed.

Critical loading conditions for the wing include +5g ultimate symmetrical maneuvering load factor, and -3.5g landing load factor in the airplane configuration with lg airload effective per Reference 11,

The weight expressions for the wing are those previously reported in Reference 12j which was part of Progress Report Number 2 for this contract, are are not repeated here. However, changes in wing peometry necessitated changes in the values of several structural constants, and the new values are -"p^orded below in Table 2.

In addition, in place of the assumption that WcA^g» the wei$it ratio of power package and propeller,is a constant, the calcu- lated values of thc-e ratios are used as they vary with disc loading and tip speed.

Table 2

Revised Structural Constants

Constant Values Used (Dimensions Are in Feet) |

1 kx 3.0

a ^.26 x 1D-2

1 P 2.222

Y .02195 x lO-2

I 6 .0182 x lO-2

1 " .02^3 x lO"2

CONFIDENTIAL «Ml» e

'

H, Feldnian 5/U/56 CHKCKCO

APPROVED

HILLER HELICOPTERS Derivation of Weight Rp Terms of Parametric Design Analysis For Propelloplane Transport Study

CONFIDENTIAL Contract Nonr 1657 (00)

JJ- 10U8

HCPOMT "c hlhS

D, Fuselage Weight static

Design arrangement and size of the fuselage is fixed by space requirements; hence the weicrht of the fuselage will vary within the range of gross weights investigated only in as much as changes in gross weight affect the general loadings.

The weight ratio for the fuselage nay be conveniently expressed by a simple analytical expression as a function of gross weight based on conventional fuselage weight prediction methods (References 13 and lU)•

E« Landing Gear Weight Ratio

Weight ratio of the landing gear is expressed by the following empirical relation (Reference 2).

«LC, .0U5 wg-02

F. Emoenr.age Weight Ratio

Empennage weight is asaumcäd to a/eraf-e 3 lb./sq. ft. of area.

Area of the horizontal and vertical tails are estimated to be (per methods of Reference 15)&

SHT - Svine ^•0217 S - -173) + l6U

.

i>VT

Defining:

bwing (•coo32 Wp'^O + ,000136 b) + .061* Ap + lh.6

W

Ap -

AR m

CONFIDENTIAL 60-OJa e

T H. Feldman j ^/u/56 MILLER HELICOPTERS 1':

APPnOVCD

r,-e: Derivation of Weight Rp Terms of Paranetric Design Analysis For Propelloplane Transport Study

loue REOOBT KIO, LlYU'^

CONFIDENTIAL Contract Nonr 165? (00)

; Total Empennage wei^iit ratio is

W er mp. 1«,

3(SVT + SHT) "< Wp AR

YT3 yr\*

.50 w .50

.000961 j JSJ* -^l1 * .000391

+ 715 .520 + .250 TZ J wH

G. Supplemental Control System

The supplemental control system consistinR of three lightweight turbo.jet engines in the aft fuselage, ducting to the tail, and a mechanism to direct the thrust in the fuselage tail cone, provides thrust required for stability and control during hover and slow speed flight.

Each engine provides 505? of the estimated required thrust so that in the event of failure of one engine, the remaining pair adequate control thrust.

From preliminary balance computations, the total thrust required is estimated to be .0U9 W^ at 6000 feet on a 95°? day.

Engine specifications, based en the information in References 16 and 17, are assumed, as follows, for sea level standard day static thrust conditions.

Specific enpine weight .110

SFC - 1.0 Ib/T/HR (normal rated power)

Idle SFC - 1.8 lb/T/HR

Ratio-idle thrust to normal rated thrust ■ .160

Ratio-normal rated S.L. thrust to thrust <i 6000 ft. 9$° ■

1.U0

Fuel for these engines is drawn from the main fuel tanks, and suf- ficient additional fuel is carried to operate two supplemental con- trol engines at normal rated S.L. power pins one (the third) engine at idle for 10 mi-utes.

CONFIDENTIAL 6(H)» e

MAMC

H. Feldnan >/V56 CHCCKCO

MILLER HELICOPTERS

APPHOVCO

T,T'-«

: Derivntion of Waight ^F Tenoa ol Psr-i.ir-Tic Desi ?rn Analysis for F: .rj-lloplane Transport Study

15

REPORT NO UoP

CONFIDENTIAL Contract Konr 16^7 (00)

Hence, the weight ratii tne throe engines» hased nn sea lev.-l^ standard day conditions, Ln terms of Te, thrust required per engine

we , .110 -S ) ■ .33 rr^ '?./

Fuel ■.-.eight ratio for two engines operating at nomal ratea power, and the third standing by at idle spe^d is

10 Te , T , L ■ ^ T/, o\ 10 »16 Te .3P1 ~

.

Ducting veight is estimated at 75 Ihs., deflecter and controls at 125 lbs., and weight of the- engine corpartrnent Including mounts, firewalls, etc. at 200 Ibs.j weight of additional fuel system, and tankare is estimated -?t 15 percent, of fuel weieht«

Thrust per engine st standard day sea level required to f'arnish the required thrust at, te;TDerature and altitude is

.P . i.u f i2ifg ^ ,0310 Wg

Hence, \Teight ratio of the entire system is

VJ scs uoo -7- ♦ .026J4

H, Wing Tilting Mechanism

Weight of the wing tilting mechanism is expressed in terms uf the study parameters by a term related to the loads of the system anci a constant representing the weight of those items which are virtaally independent of loading chanrres within the range of rross weightö investigated.

Investigation shows good correlation is obtained from the following expression.

WQI = l.p'5 Wj + 200

Wj is the weight of both jack shafts and is related to the loads and reonetrv as follows:

CONFIDENTIAL 6(H)» C

H. Feldnan CHCOKCO

AF^KOVCO

OATC '"'

5A/56 HILLER HELICOPTERS TIT'-=:Derivrtion of Weight RF Terms of

Parametric Desifm Analysis for Propelloplane Transoort üzndy

lOhB »C^ORT KO. U7U.5

CONFIDENTIAL Contract Nonr 16^ (00)

Critical loadinp; condition oocurs durinr a suddenly braked landinn; roll viith winft tilted to vertical nosition. Horizontal load factor M3 = -1.75 ultimate.

The shafts of effective column iength .55 Croob resist a load resulting from l.T^Cwing vieight + pc.'jer package weight + propellei weight + fuel and fuel tank weight) appiied at the err- of the hinged mass. Total column load, from system geometry, is

P = 2.66 \We

jpp Wt h r?— + ^r.

Wj

vT VJ FT V;g

-e

••.

and assuming each shaft good ?or 3/u of total load to nrcvide for asymmetry of loading, design loading for one shaft is

cr ■i x/vg mg "g wg rtg

Wj, weight of two threaded hollow steel shafts is approximated"-

Wj - 2(1.5) 5; (Do - Di ) [-55 (l2)CRn .3 - ä.57 ^ Do - Di^) CR

sc

W-

12j hence Di = .^35 D0 ; Croot = 1.5 Gavr. - 1. C

2.12 C Do'

Do is defined in terms of Pcr on a long column

n 2 El or i2~ (^(l^) .55C]2 (10) 2pWf

Ireq'd " 6*65C2 ri'.rc (13)"7 - -^ (D0

Ü - D^)

Do2 - .0051L C pi Wgi

CONFIDENTIAL 60-ox e

H. Feldnan CHCOKCD

-■

Sßtäk i MILLER HELICOPTERS

CONFIDENTIAL

Tin.«. Derivation of Weight Hp Terns of Parametric Desigr. Analysis for Pronclloplane Transnoi't Study

10U8

RCPOItT NO. U7Uo5

Contract Monr 1657 (00)

Her.ca,

W, 2,12C (.0051a Cp: Wgi) .0109 C'p'^ VJg2

and

H/K ,0202 C 2 1

^L

200

«8 " V

Notinrr vrinpr chord,

c . r we 1 2

A a

Wg

Wvm .0202 W

^ (1) 2'j0

In calculating p, the fuel and fuel tank weight is necessarily assumed for first approximation

hence. «H , Wpp , Wp vT + u wT + '4 w: + 3L#l5 RF

I. Fixed and Operational Equipment

Weight ratio of fixed and operational equipment is estimated in the conventional manner and is tue ratio uith respect to .rross weight of the following items.

1. Propeller Equipment Weight (Includes weight of spinners, anti-icin^ eq'iipment, synchronizer, electrical beta and ey.tc controls, etc.) 800 lbs.

2. Air Conditioning and Anti-icing Equipment ^OO Ins.

3. Electrical System (Empirical). , . , . . .00375 Wp * 7!>0 lbs.

U. Instruments and Navigation Equipment . 500 lbs.

$. Electronics ........ 500 lbs,

6. Hydraulic and Pneumatic Systems ..... .005 Wp + 150 lbs.

CONFIDENTIAL «Mm e

H. Feldman CHCCKCD

5/ii/$6

AP^HOVCD

HILLER HELICOPTERS TITLE: Derivation of Welpjiit Hp Terris~c

Parametric Desifpi Analysis for Prepelloplane Transport Study

lOhQ »CPORT -o. ii7J4,5

CONFIDENTIAL Contract Nonr 16^7 (00)

7 c Surface Controls .•••«••...•« «010 W-r + c')Q Lbso

8, Misc. Fiornishines, Acconodations, Equipment . „ , 20^0 lbs.

Hence,

WpCE ^300 vZ + .02375

For a 70,000 lb. aircraft, the fixed and operational equipment hence weighs approximately 7,000 lbs. exclusive of win^ tilt mechanism and supplemental control system.

CONFIDENTIAL 60-030 e

I

PRCPARCO

CHECKED

APPROVED

NAME

H. Feldiian DATE HILLER HELICOPTERS

TITLE Derivation of Weifrht ^tp Ferms of P?.r.'inetric Desiqn Ar.alrsis for Pronello^lane Transport Study

C 0 N F IDENTIAL Contract i-cnr r.),"" (Of)1)

riQORE 1.

PROPFXLLit IxP DiSI^' MOISNTS

lOiifi

NO. U714.5

MQ «) Zero iJt,ation vs. Disc Loading«

V

.05

•Oit

(Single Rotation eo3 Propellers)

.02

.01

0

.06

.05

.Olj

.03 *)

(Dual Rotation Propellers) ^02

.01

;i0

aO

■« 3,

Transition Critical

Flight Conditions Critical

6o ao wH (lb/ft2)

100

6J o0 wH (lb/ft2)

1D0

CO

Mo*.Mo/|(^-)^

^transition " [<Cy ^ ^ ^ '" + ^ TV75 ^ -

"Ofli-rht " M W Ai•^ D3A?)co3 p,7 (n.^ ♦ 2CL,7 cot j-17)

TTTTTTTTTTT"-"

PRCPARFDNAME

L-, Feldr;an ^/n/% HILLER HELICOPTERS fag*

CHECKED TiTH:yiori_-03ticn of /.’eit'ht Hj? Terr.? of Psrarietric Lesip;, Analysis for Pro ' ellonlane Transr.ort Study

MODfL

APPROVED REPORT NO.

CONFIDENTIAL Oontraot Nonr 165? (X')

i a.' : (i fcij 2 »

liATIG OF PL. OT PiOPS VO. 3(A?)V7 « ^'00 FP3 Go,yyj;/

ell

*10

.09

-0?

.07

Wpp +V)p200

One Nacelle Only

i;OT' 60C’B(A?)

rO(! 1000

B(AF)LEG:iND

jlaxinuj!i Allcn-rable Otall for Control Dnrir.p: liover-

4® otractural Limit 3 or 6 blades, structural Limit h or P- blades.

9 OptirauiTi r (Ar) .CONFIDENTIAL

«0428

MAMC

H, Fsld!;;ar.OATK

5/L/56 HILLER HELICOPTERS «ac 21

CHceicco TiTu; Derivation cf Weight Rf Terms of Parametric Design Analysis for Pro.elloplane Transport Study

MOOCL 10,1,8

A^^NOVCO

---------------------- 1

!Gor.tract Nonr l657 (00)

.GFGREKOiLS

1. Tra.isnort Prooelloolsuie Stuch'- Prorress Report No. h for Contract Nonr 1657 (oo), Hiller Helicopters, December, 1955.

2» Helicopter Propulsion System Stud;/- for USAF Contract AF 33(03fi)22185, Thermal Research and Engineering Corporation, September,1952.

3. Transport Helicopter Design Anal^^-sis Methods for Contract Nonr 13i»0 (00), Hiller helicopters, November, 1955.

li. "An Approximate Propeller Vleii^t - Strength Relationship",Curtiss-tJright Corooration, Propeller Division, ICH U777-E,October, 1955.

5. "Investigation of the Aerodynamic Characteristics of a Model Wing-Propeller Combination and of the «ing and Propeller Separately at Angles-of-Attack up to 90 Degrees", NACA T.N. 330h, November,195li.

6. "Aerod:manic Investigation of a Foiu* '-lade Rropoll:r Operating Through an Angle-of-Attack Ran'-s from 0*^ to -I'O^ ", l:ACA T.N.3228, June, 195U.

7. "A Simplified Theoretical Investigation of a ling-Rropeller Combination Through a Ran'-c of Awgles-of-Attac': From 0 to 90 and a Comparison with Experimental lesults", nhLller Helicopters Report No. 161.3, Octo'^er, 19'-'o.

8. "A Method of Estimating the Propeller Shaft Loaho nd iced by First- Order Propeller Excitations", Crrtiss.^iright Corporation, Proneller Division, Report No. C-2378, August, 1952.

9. "The IxP Propeller Vibration Problem and Related Effects", Curtiss- Wri ht CoipDoration, Propeller Division, Report No. C-2131, January, 1950.

10. "Procedure and Data for Propeller Performance Analysis", Aeroproducts Division Report No. 190, April, 19iUi.

11. "A Statistical Study of Wing Lift at Oround Contact For Four Trans­

port Airplanes", NACA T.N. 3h25, April, 1955.

60-030 C

PmtfARCD NAMK

H, Feldman CHICKED

| APPBQVCO

S/h/S6 HILLER HELICOPTERS -•<.• .-: T,TLt: Derivation of Weight Hp Terms of I MOO,t

Parametric Desipn Analysis for | — Propelloplane Transport Study | '*">•••'-Q

(Contract Nonr 16^7 (00)

:o2»8

i ;:.

REFERENCES

12. "Variation of Propelloplane Wing Weight with Ciross Weight, Wing Loading, and Aspect Äatio", Hiller Helicopters Renort No, U61.1, July, 1955o

13. Hopton-Jones, P., "A Practical Approach to the Problem of Structural Weight Estimation for Preliminary Design", National Society of Aeronautical Weight Engineers, May, 1955.

111. Corning, G., "Airplane Design", Edwards Brothers, Inc., 1953«

15. Perkins, C.D., and Hage, R.E,, "Airplane Performance, Stability and Control", John Wiley and Sons, Inc., 19i49«

16. "Model Specification, Xj8l^fE-3 Turbojet Aircraft Engine", Westinghouse Aviation Gas Turbine Division, WAGT-216B-B, January, 1955«

17. "Rolls Royce Soar R.Sv. 2 Initial Project Performance Data and Installation Notes", TSD Publication h97, September, 1953.

60-030 C

I

UNCLASSIFIED

UNCLASSIFIED