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Produced by the NASA Center for Aerospace Information (CASI)

NASA Technical Memorandum 78715

(NAS - T" -7871 5) EPPECTS 0 L MINAR fLO ~ N79- 178 1 CO TROL HE E A CE f A LARGE SPAN-DISTRIBUTED-LOAD LYING-WIN CARGO A Rpt I'TE CO CEPT IN SA ) 23 P H A0 2/ F 0 1 nclas

C CL 0 1C G /05 1 19 1

EFFECTS OF LAMINAR FLOW CONTROL ON THE PERFORMANCE OF A LARGE SPAN-DISTRIBUTED-LOAD FLYING-WING CARGO AIRPLANE CONCEPT

LLOYD So JERNELL

JUNE 1978

NI\SI\ NaTion I Aeronaullcs and S AdmlnlSTrallon

t.ng!oy Research Conler Hampion Vorglnla 23665

\

, I

\

/' I

~FFECTS OF LAMINAR FLOW CONTROL ON THE PERFORMANCE OF A LARGE SPAN -OJ STRJ BUTED-LO.~I) FLY ING -WJ NG CARGO AIRPLANE CONCEPT

by Lloyd S. Jernell

Langley Resea rch Center Hampton, Virginia 23665

SUMMARY

A study was conducted to determine the effects of laminar flow control on the performance of a large span-di s tr'i buted- l oad flyi ng -wi ng cargo ai rp lane concept having a design payload of 2.669 MN (600 000 lbf ) and range of 5.93 Mm (3 200 n.mi.). Two configurations \~ere consi dered. One emp loyed laminarized flow over the entire surfaces of the wing and vertical tail s , with the exception of the estimated areas of inte rference due to the fuselage and engines . The other case differed only in that laminar flow was not applied to the flaps , e levons , spoilers , or rudders. The two cases are referred to as the 100 percent and 80 percent laminar configurations, respect ively.

The ut i l ization of l aminar flow control results i n reductions in the standa rd day , sea level ins talled max imum sta ti c t hrust per engine from 240 kN (54 000 lbf) for the non-LFC confi guration to 205 kN (46 000 lbf ) fo r t he 100 per~ent laminar confi guration and 209 kN (47 000 lbf ) for the 80 pe rcent case. Weight increases due to the LFC systems cause increases in t he ope ra ting empty we ights of approxi mately 3 to 4 percent . The des ign takeoff gross weights dec rease approximatel y 3 to 5 percent. The FAR-25 takeoff field distances for the LFC configurations are greater by about 6 to 7 pe rcent . Block times are virtua1ly unaffected by the util ization of LFC. As compa red to the non-I.FC configuration, block fuel weights are reduced 24 percent for the 100 l a~inar conflgurat ion and 18 percent t ~ r the 80 pe rcent case. Fue l efficiencies for the respective configurations are i nc reased 33 pe rcent and 23 percent.

INTRODUCTION

The endeavor of providing the most effi cient ai rpl ane practi cab le has become es pecially important in recent years beca use of the increasi ngly high cost of aircraft fuels. One method of imp roving ae rodynami c effi ciency is that of reducing drag through laminar flow control. In this thod, suction is employed to remove a portion of the bounda ry lJyer through sma ll pe rforat i ons or slots in the aircraft sk in . The effect is to shift the puint of transition downs t ream, thus increasi ng the area affe cted by laminar flow.

Based on he ~ss umption tha t a prac i ca 1 laminar flow control sys tem may be available in the 1990' s , a s tudy was conducted wherein the effects of a laminar flow control sys tem on aircraft performance were evaluated for th air­plane concept of reference 1. This aircraft is envi sioned as a type of large cargo airplane possibly becoming operational in about the same time period. A configuration study (ref. 2) was performed by the Vought Co rporation, Hampton Technical Cen ter, under the auspice of the Vehicle Integration Branch , Aeronautical Sys tems Di vision , Langley Research Cen te~.

The purpose of the present pap r is to summarize the con tracted analysis and , by compar ison with the st udy of reference 1, to indicate the gains whi ch cou ld be expected from the application of laminar flow control to an advan ced f ly ing-wing ca rgo transport.

SYMBOLS AND NOTATION

Values are presented in both 51 and U.S . Customary Units . The measurements were made in U. S. Cus tOl'1ary Un i t s.

c

CD

C D,Pmin

Cf

CL

rp

CSP

DOC

f

KEAS

LFC

M

OWE

q

R

2

local chord

mean

drag

aerodynami c chord

coe ff icient Ora , qS

mill imum pa ra si te drag coe ffici en t

total avera ge sk in fri ction co ffi cient

lif t coeffi cient, Lift q

pressure coe fficient

suction power coe ffi cient , Sucti on power 12PO(N~S

di rect ope ra ti ng cos t

rq uiva1 nt flap plate area

eq uival ent ~i rspeed . ~no s

laminar f low control

~la ch number

operating wei ght. empty

dynami c press ure

Reynolds numbe r

S reference wing area

Sv area per verti ca 1 tail

s surface dis tance along airfoil. measured from leading edge

T sea level. standard day installed stati c thrust per engine

TOGW takeoff gross weight

TSFC thrust specific fuel consumption

V velocity

Wf fuel weight

p densi ty

Subsc ripts :

local

1 am 1 ami nar fl ow

s suction slot

turb t urbulen f low

freestream

DISCUSS I Otl

Configu rations

The study was conducted to determine the effects of the applicat ion of lami nar flow control (LFC) on the performance of the large f lying-wing cargo a irplane concept of reference 1. (See figure 1 for configuration detail s). The missi on requirement of the referen ce aircraft wa s that of transporting 2. 669 MN (6DO 000 lbf ) of container·zed cargo (maximum container cross section of 2.44m x 2.44m (8 ft x 8 ft ) over a di stance of app roximately 5.93 Mm (3200 n.mi. ) . The primary design philosophy in configuring the aircraft was to distribute the payload along the wing span in order to counterbalance the aerodynami c loads as much as poss ible. thus minimizing the in-fli ght wing bend ing moments and shea r forces.

In the LFC s tudy the ex terior dimensions . exc ludi ng nace ll e size . remained unchanged since these dimensions are determined almost soley by the cargo di men­sions . Engine type and design payl oad and range were he ld constant. Engine si ze . ope rating empty weight. and mission fuel wei ht were adjusted to those required for the LFC configurations . Although the locations of the suction pumps were not

3

consi dered in the study , ample unused wing volume rea rward of the wing rear spa r is available for the pumping systems. Several independent systems would be used to minimize the length of ducting between the slots and the pumps.

Two LFC configurations were evaluated. The first employed an LFC system over the entire surfaces of the wing and vertical tail ~ , with the exception of t he estimated areas of interference due to the fuselage and en9ines. These inte rference areas were 37.2 mZ (400 ft 2) and 24.5 m2 (264 ft 2) for the fuselage and engi nes, respectively; or only approximately 1. 5 percent of the tot~l wetted area of the wing and vertical tails. The second configuration differed only in that the LFC system was not applied to the flaps, elevons, spoile rs, or rudders . These t wo configurations are referred to as the 100 percent and 80 pe rcent laminar cases, respectively.

LFC System Weights

The wing wa s assumed to be of aluminum honeycomb construction . The LFC sys tem weights for the two configurations are presented in table I. The values of weight increment per unit surface area were obtained from progress reports submitted during the course of a NASA-contracted laminar-flow-control study currently being conducted by the Boeing Comme rcial Airplane Company. It should be noted that these weight increments include the weight of the associated pump­ing and ducting sys tem, as well as the incremental weight of the surface structure.

Drag Polars

The drag polars for the 100 percent and 80 pe rcent laminar configurations are shown in figure s 2(a) and 2(b) along with the polars representing the non-LFC case (ref. 1). It was estimated that there woul d be a one-th ird reduction in interference drag at the j unc ture of the wing and vertical tails because of flow laminarization . Typica l calculations of the min i mum parasite drag coefficients, at a ~ruise Mach number of 0 .75, are presented in table II.

A maximum Reynolds number of 6. 56 x 106 pe r m (2 .00 x 106 per ft) wa s specified for the effective ope ration of the LFC system . This Reynolds number limitation required a reduction in the r limb velocity from 280 KEAS for the non-LFC configuration (ref. 1) to 250 KEAS. Performance claculations (subse­quently to be discussed) indicate that the LfC system shoul d be acti vated when the aircraft reaches a Mach number of 0.72 and an alti tude of 9.69 km (31 800 ft). The cruise Mach number of 0.75 i s the same as that for the con­figuration of reference 1. Hence, the LFC dra9 po l ars are app licable only during the latter part of the climb ( M ~ 0. 75) and throughou the crui se segment .

4

. ~-

LFC Power Requir ment and Equ ival ent Drag

At a given location on the surface of an LFC configuration the coefficient of suction power required to operate the sys t em (exc luding losses due to ducts , valves , and pumps) may be approximated by the eq uat ion

( 1 \

Equation 1 is essentially the same as that de r ived in reference 3, except that the local and freestream densities are not as sumed to be equa l. The airfoil pres su re coefficients, calculated by the method of reference 4, are shown in figure 3 for the de s i gn Mach number and 11ft coeffi cient. The sucti on f1 ow velocity, Vs ' and density, " were determined with the use of the computer program described in reference 5. This program calculates the compressible boundary layer and suction flow characteri s tics over a yawed infin ite wing. Constant velocity suction flow along segments of the chord normal to the lead­ing edge were input into the program in a tria1 -and-error manne r until complete 1aminarization wa s achi eved . For the present study, the finn1 di stribution of the ratio of local suction flow dens ity to that of the freestream density is shown in figure 4. The variation of the ratio of suction flow velocity to that of the freestream velocity is presented in figure 5. The resu ltant distribution of the suction power coeff ici ent along the chord, as determined by equation 1, is shown in figure 6. These data were integrated to determine the sucti on pow r coeffi cient s of the 1aminari zed spanwi se stations of the airfoils . The integ ra­tions were performed to the trailing edgf for the 100 pe rcent laminar case , and to the contro l-s urface hinge lines for t:,e 80 percen t case . Si nce the wing is untapered, the suction power coeffi cient is invariant along the span, and thus, t he suction power coefficient of the wing i s eq ual to that of the ai rfoil except for a reduction to account for the non 1aminari zed areas. This reduction was es timated to be 10 pe rcent. Th us

eSP ,wing = 0.9 CSP ,airfoil (2 )

The values of suction power coefficient fo r the wing are 0.0015 and 0.0014 for the 100 pe rcent and 80 percent laminar cases , respectively. The suction power coefficient for the tails could not be ca l cul ated by the same n~thod sinr.e the airloads were not as well defined as those for the wi ng . It wa s therefore ass umed that the magnit ude of the suction power coeff icient (based on the refer­enced wing area ) for the tails relative to the wi ng is proportional to the areas of the respective components. The refore , the total suction power coe ffi ci ent for the a i rp 1 ane is

CSP , total = 0.9 CSP ,airfoil (1 + 2 ~v ) (3 )

5

The total suction power coeffi cients are 0.0017 and 0.0016 for the 100 percent and 80 percen t laminar cases , respectively .

The suction power may be expressed as

suction power = CSP( ~",V~) (4 )

= ( CSp t"'V&S ) V", ( 5)

Since power equals the product of force times velocity, the te~ within the brackets of equation 5 may be considered an equivalent drag, with CSP an equiv­alent drag coefficient. Thus , for each LFC configuration , the equiva lent increase in drag coefficient due to the total suction power requirement is equal to the total suction power coefficient determined by equation 3.

Weight Adjustments, Engine Resizing , and Performance

As previously mentioned, the aircraft exterior dimensions , exc l uding those of the nacelles , remained unchanged, as did engine type , and design pay load (2. 669 MN (600 000 lbf)) and range (5 926 km (3200 n.mi . )). How,ver , engine s ize, operating empty we ight , and mis sion fuel weight were adjusted to the requirements for each 0 " the LFC c ~"' figurations. Mis s ion performance was evalu .. ated with the use of the Vehi cle Integration 8ranch long-range-crui se mission anal ysis program develo~ed at the Langley Research Center. The fuel requiremen ts for ta xi, takeoff, and des cent were adjus ted to refl ect changes ina i rcraft wei9h t and engine size.

The required operating empty weight, mis sion fuel weight, and engine size for each LFC configuration were determined in two basi c steps. In the first s tep t he desi gn range and engi rle size were ass umed t o be eq ual t o those of referenCE' 1. Us ing the mi ss ion analys i s program, fuel weight and ope rat ing empty weight (as a funct i on of fue 1 wei ght ) were determi ned. Fo r the 100-pe rcent-LFC confi gurati on, the ope rating empty weight was 1. 872 MN (420 800 l bf) and the mission fuel weiqht was 1.359 MN (305 500 lbf ). For the 80 pe rcen t LFC case t he ope rating empty weight W3S 1.846 MN (4 14 900 lbf ) and the fuel weight was 1.438 MN (323 300 lbf ) .

In the second s tep these mi ss ion fuel wei ghts were hel d constant and the operating emoty weights were varied as a funct 'ion of engine size (or thrust) , These weight s were then input into the mis sion analysis program to determine range and performance. The estimated effects of engine thrust on operat ing empty w~ight, the resultant design takeoff gross weight, and the calculated range for the 100 percent and 80 percent laminar con f i gurati ons are shown n figures 7( a) and 7(b ) , respectively . For engine th rus t s less than those shown in these f i gures , the servi ce ceilings at the begi nni ng of crui se were too low to meet the aforement i oned maxi mum Reynolds number r'equi remen t for effec ti ve LFC operati on. The to t al dec rease in the weight s of the si x engines are 259 kN (58 300 lbf ) for the 100 pe rcent LFC case and 223 kN (50 200 l bf ) for t he 80 percent confi gurat i on. The mis sion pe rforman ce of t,le 100 pe rce nt and

6

80 percent laminar confi9uration~ with th minimum-scale engines are presented in tables III(a) and III(b), respecti vely .

Effects of LFC

The more si gnificant parameters for the two configurations studied are compared with the non-LFC configuration (ref. 1) in table IV . The reductions of both drag and gross weight at the b ginning of cruise (due to the application of LFC) results in decreases in the sea level, standard day installed maximum static thrust per engine from 240 kN (54 000 1bf) for the non-LFC configuration to 205 kN (46 000 lbf) for the 100 percent laminar configuration and 209 kN (47 000 lbf ) for the 80 percpnt case.

Although there are small decreases in the weights of some of the wing primary structural components of the LFC configurations due to lower fuel weights , the additional weights due to the LFC systems result in small overall increases in the operating empty weights of 4.1 percent and 3.3 percent for the 100 percent and 80 percent laminar configurations , respectively.

Because of the significant reductions in cruise drag afforded by the LFC systems, mission fuel weights (block fuel plus reserves minus taxi-in) are reduced 21.8 percent for the 100 percent LFC configuration and 16.2 percent for the 80 percent laminar case. The resulting des;gn takeoff gross weights (operating empty weight plus mis~ion fuel weight plus the 2.669 MN (600 000 1bf ) payload) decrease 4.8 percent and 3.5 percent for the 100 percent and 3.5 percent for the 100 percent and 80 percent LFC configurations , resp ctive ly .

The FAR-25 takeoff field distance, whi ch is 2.5 km (8200 ft) for the non-LFC con figuration, is increased by approximately 159 m (520 ft) for the 100 percent laminar configuration and by about 177 m (580 ft) for the 80 percent case. Although takeoff field distance would be shortened for the laminar configurations due to the lower wing loadings, thi s advantage is slight ly outweighed by the lower thrust-to-weight ratios.

Block time (engi ne start to engine shutdown ) is virtually unchanged from that of the reference 1 configuration. As compared to the non-LFC configuration, block fuel weights are reduced 24 percent for the 100 percent laminar configura­tion and 18 percent for the 80 percent case. The fuel eff iciencies for the respective configurations are increased 33 percent and 23 percent.

Lack of operational experience with laminar-flow-control aircraft precludes an accurate est imate of airframe price or maintenance cost. Therefore, direct operating costs were estimated assuming that all DOC parameters except fuel are equal to those of the reference 1 configuration. For a fleet size of 100 air­craft and a fuel price of $0 .42 per gallon, the reduction in DOC due to the application of LFC is 10 percent for the 100 percent LFC con figuration and 7 percent for the 80 percent case. For t he same fleet size and a fuel pri ce of $1.20 per gallon, DOC is reduced 20 percent for the 100 percent LFC configurat ion and 17 percent for the 80 percent case.

7

CONCLUSIONS

A study was conducted to determine the effects of laminar flow control on the performance of a large span-distributed-10ad flying wing cargo airplane concept having a design payload of 2.669 ~N {600 000 1fb} and range of 5.93 Mm {3 200 n.mi.}. Two configurations were considered. One employed 1aminarized flow over the entire surfaces of the wing and vertical tails, with the exception of the estimated areas of interference due to the fuselage and engi nes. The other case differed only in that laminar flow was not applied to the flaps, e1evons, spoilers. or rudders. The two cases dre referred to as the 100 percent ar.~ 80 percent laminar configurations, respectively. The conc lusions are as follows:

1. The utilization of laminar flow control results in reductions in the standard dar. sea level installed maximum static thrust per engine from 240 k;j (54 000 1bf) for the non-LFC configuration to 205 kN {46 000 1bf} for the 100 percent laminar configuration and 209 kN {47 000 1bf} for the 80 percent case. ,

2. Weight increases due to the LFC systems cause increases in the operating empty weights of approximately 3 to 4 percent. The design takeoff gross weights decr~ase approximately 3 to 5 percent.

3. The FAR-25 takeoff field distdnces for the LFC configurations are greater ~y about 6 to 7 percent.

4. Bloc', times are virtually unaffected by the utilization of LF C.

5. As compared to the non-LFC configuration, block fuel wei ghts are reduced 24 percent for the 100 laminar co nfiguration and 18 percent for the 80 percent case. Fuel efficiencies for the respective co nfig~rations are increased 33 percent and 23 percent.

REFERENCES

1. Jernell. Lloyd S.: Preliminary DeSign of a Large Span-Distributed-Load Flying-Wing Cargo Airplane Concept. NASA TP 11 58 , 1978.

2. Lovell, W. A.; Price. J. L; Quartero, C. B.; Turriziani. R. V.; and Wa shburn. G. F.: . Design of a Large Spa n-Dist.ributed Load Flyi ng-Wing Cargo Airplane With Laminar Flow Control. NA SA CR-145376. June 1978.

3. Ri ege1s . Friedri ch W.: Airfoil Sections. Translated from German by D. G. Randall. Butterworths, London, 1961.

4. Bauer , F. ; Garabedian, P.; Korn , D.; and Jameson, A.: Supercritical Wing Sections II. Lecture Notes in Economics and Mathemati cal Systems . M. Beckmann and H. P. Kunzi. eds .• Springer-Verlag , 1975.

5. Carter . James E. : A Fortran Program for the Suction Transition Analysis of· a Yawed Wing Laminar Boundary Layer. NASA TM X-74D13 , 1977.

· ~

TABLE I. - LFC SYSTEM WE IGHTS

100 percent lami nar ; zed conflgurat ion 80 percent l amlnarlzed conf lqurat ion

Proj ected Weight Weight Pr oje<:ted Weight Ve igh t

Lamlnarized ; ncrernent i nc ren'lent laminarized increonent increaent

area. nl2 (ft2) per un; t a re~ I N (1 bf) area • .,2 ( ft Z) per uni t area I N ( Ibf) N/.,Z (I bf/ ft )

Wing (s tructure , duc ts . 1, 662 . 6 ( 17 ,896) 60 . 329 ( 1.26) 100 , 303 (ZZ ,549) 1 ,310.7 (1 4,108) 60.329 (1. Z6 ) 79,073 (17 ,776 ) and va l ves)

Suction eng; ne for 1,652.6 (17 ,896 ) 33 .516 (0.7) 55 , 724 ( 1 Z , 5Z7) 1.310.7 (1 4 ,108) 33.516 (0.7) 43 ,929 (9 ,876) wing

Suction e~ine for 41 9. 4 (2 , 362) 33 . 516 (0.7) 7 ,353 (1 ,653) 188.8 (2 .032) 33.516 (0.7) 6,328 (1 ,423 ) vert ical ta il (assumed to be part of wi ng weight ,

Tota 1 i "crease in 163 ,380 (36,729) 129 ,330 (29 ,075) wing weight due to LFC

Ver t iea 1 ta il5 219 .4 (2,362) 60.329 (1.26) 13 , Z36 (2 ,976 ) 188.8 (2 ,032 ) 60.329 (1.26) 11 , 390 (2 ,560)

To ta l OWE Incr ease 176,616 (39, 705) 140,720 (31 ,635 ) due t o LFC

- ----

TABLE II. - CALCULAT ION OF MINIMUM PARASITE DRAG COEFFICIENT. ~ • 0 . 75

Airplane R ~rag cJmponent x 10-1 Item

Wing 118 .( Uncorrected flat plate

Superveloci ty

Pressure

Roughness

Excrescencies

Wing-body interference , ,. To ta l . w.ng

Ta i1 s 44.6 Uncorrected f lat plate

Supervel ocity

Pressure

Roughness

Excrescenc i es

Wing-tail interference , ,Ir To ta l . tail s

Total . wing and tail s

To ta l ai r plane Co ' I'm; n

-- -- - ------- ---

Total turbulent CD ,Pmin

0.01059

l urbulent l Uln" lam.nar

Cf f Cf 2 ft 2 ,,,2 m

0.001921 6 .994 75.28 0.00011 0 0. 400

2.239 24.10 .128

.670 7.21 .038

1 . 211 2.27 . 012

.622 6.70 . 311

. 186 2 .00 .186

10 .922 117.56 ,Ir

1.075

.002210 0.980 :J .55 .000180 0.080

! . I .151 1.62 .012

.005 .05 .001

I .031 .33 .003

I .085 .91 .042

~ .931 10.02 .621

2.183 23. 48 ,

.759

13 . 105141.04 1.834

• 0.00405

* Total laminar C!) .~ . m.n

If - f I " 0.0105~ _ ' turb lam wing. tail s

s

!IO'.I 10. no r f

ft2 Cf .2 ft 2

4 .31 0.000546 2.012 21.66 , 1.38

I .644 6.93

.41 .193 2.08

.13 .012 .13

3.35 .311 3.35

2.00 .186 2.00

11 .58 ~ 3.358 36.15

.86 .000505 0.22C 2:l

.13 I .03C .37

.01 I .001 .01

.03 I .003 •03l

.45 .042 .45 '

6.68 I .621 6.68

8.16 ,

.925 9.9' l

19.74 I L-. ...,

4.283 46.10

• 0 .00547 I

TABLE III . - MISSION PERfORMANCE

(a ) 100 percen t LFC Configuration (Taxi - In fuel taken out of reserves at des tination .

Clv l ' A.rondutl cs Board range equal s trip range on lr. . allowances for maneuver , traffic, . nd airway distance)

(a) Aircraft characteris t ics

Takeofr gros s wei ght, N (Ibf) ... . ........... .. .. . Operating we igh t, emp t¥ , N (Ibf) . ... ... . .... •. ... Payl oad, gro~s , N (I bfl ........................ .. Wing area , nI ((t2) .... .... ............. . .. ..... . Installec sea- l evel s tati c thrust, pe r engine ,

standard day , N( I bf) .......................... .

5 760 44 7 I 790 409 2 668 933

I 724

204 618 Takeoff thrus t -we igh t rat~o .... ' 2'" ...... . .. . . . Takeoff wi ng loading, N/m (Ib f/ ft ) ........ . . . . . 3337

(b) Oesl9n missi on

F1 Igh t Mode Gross . el9ht , N (Ibf)

6Fuel , N ( Ibf )

Takeoff ........ . 5 760 44 7 ( I 295 000) 21 351 (4 800)

Start climb 5 739 096 ( I 290 200) 152 574 (34 300)

Start cruise 5 586 522 (I 255 900) 826 480 185 800

End crui se 4 760 04 2 (I 070 100) 19 127 (4 300)

End descent 4 740 915 ( I 065800)

(I 295 000 ) (405 SOOl (600 000 (18560)

(46 000 ) 0. 213

(69.7)

ARange , km (n.ml.)

o

393 (2 12)

5154 (2783 )

370 (200)

'Ti me mi n

II

36

386

20

Taxi-In ...... . .. 4 734 087 ( I 064 265) 6 828 (1535) ___ -"0 5

Block fuel and time................... 026 360 (230 735) Trip range . .. . ... ... . • .. ..... . . .. ............ .. .. .. .. . . . .. 5917 (3 195)

(c) Reserve fuel breakdown

10-percent tri p t ime , N (Ibf) .. .. ........ .. .. . ... . Mi ss ec approach, N ( Ibf) ........................ .. 370 km (200 n. mi.) to alternate airport , N (Ibf) .. 30 minutes holding at 457 m (1500 ft), N (Ib f) .... lota 1 reserve fu e l

87 630 (19 700 ) 15 124 (3 400)

11 3 874 (25 600) 64 944 (1 4 600)

281 572 (63 300)

(d) Initial crui se conditions

CL

... .... ............. ... .. ....... ... ..... ... .. .. .

Co .. . .. ... . . . .. .. .... ... .. . .....• . .• , •. •... , .• .. . , LID . .... . .. .... . .. .. .... . ........... , .. •. . ........ TSFC , kg / N-hr (Ibm/lb f-h r ) .............. .. .... , .. . Alti tude, m ( ft ) ................................. .

(0 ) rucl efficiency

Pay load of fucl bu rnec, Mg-km/ N ( ton -n. ni ./lbf)

0.3207 0.01253

25. 60 (0.636)

10211 (33 500)

1.57 (4.1 5)

458

TABLE III . - MI SSION PERrORMANCE (b) BO percen t LFC configurati on

(Taxi-in fuel uken ou t M roseryes at destination . Ci yil Aeronaut ics Board range oqu.ls tr ip range mi nus al lowances for maneuver. traffic and airway dlst.nce)

(a) Aircraft Character is ti cs

TJkeJff gross weight . N (lbf/ .......................... .. S 839 625 (I 3(2800) Operating we ight. empty. N ( bf) ....................... .. Payload. grop. N (Ibf) ................................ .. Wi ng area, m (f t 2) .................................... . Installed sea- l evel s tat ic thru st , per engine, s tandard

I 777 065 2 668 933

I 724

day. M, Ibf ......... . . . .......... . ...... . ............. . 209 066 T,.keoff thrust-weight rat~o 2 Takeoff wing loading. N/m (Ibf/ft) .. ......•..•.....•... 3385

(b) Desl9n missio n "Range .

Fli ht Mode Gross wel9ht, N (Ibf) tJue 1 I N (Ibf) km (n.ml.)

Takeoff ........ 5 839 625 (I 312 800) 21 633 (4 870) 0

Su rt cl1 .. b . .•• 5 817 963 ( I 307 930) 160 580 (36 100) 419 il26) Start cru ise ... 5 657 382 (I 271 830)

902 099 (202 800) 5137 (2774)

End cruis e .... 4 755 282 (I 069 030) 19 572 (4 400)

End descen' 4 735 710 (I 064 630) Ta x i-in 4 726 726 (I 063 060) 6 964 (I 570) Blcok , •• 1 and time .................. 110 899 (249 740 )

Tri p range

(c) Reserve fuel breakdown

10-percent tri p ti me . N (Ibf) ............ 95059 (21 370l Missed approach . N (Ibf) ................. 15 435 (3 47 0 370 km (200 n.mi.) to alterna te airport .. 114 230 (25 680 30 mi nutes holdin9 at 457 m (1500 ft)

N (I bf) ........ .. ...................... 65 166 (1 4 650) Total r~serye fuel ...•......•......•..... 289891 (65 170

(d) Initial cru ise conditions

CL. .............................................. 0.3404

Co ... . .... . ..... .. .. ..... . ..... ..... .......... . . 0.01462

LID ............................. ................ 23.28

TSFC, kg/N-hr (Ibm/lbf-hr) ... .. .•. . •... .. ... .. . . (0 .628) Altitude , m (ft) .......................... 10516 (34 500)

(e) Fuel ef' (ciency

Payload of fuel burned. Mg- km/ N (ton-n. mLll bf) ............................ 1. 45 (3.84)

370 (200)

0

5926 (3200)

!399 SOOl 600 000 (18 560

(47 000) 0.215

(70.7 )

~Time . mi n

11

37

386

20

5

459

· ),

Thrust per engine, N (lbf )

Operating empty weight, N (1 bf)

Mission fuel, N (lbf) (Block + reserves - taxi-in)

Design TOGW, N (1 b f )

Takeoff field length, m ( ft )

Block ti me , min

Block fuel , N (lbf)

Fuel efficiency , Tonne-km per N of fuel ( ton-n .mi . per lbf

('f fuel)

TABLE IV. - EFFECTS OF LFC

Design range ' 5926 km(3200 n. mi . ) Design payload = 2 668 933 N (600 000 lbf )

Non-LFC 100% 1 ami nar r' changeJ from non~

Confi gUl j ~ ion confi guration LFC conf

240 200 (54 000) 204 600 (46 000 -14.8

1 719 682 (386 600) 1 790 409 (402 500 4 .1

1 663 635 (374 000) 1 301 105 (792 500 - 21.8

t 052 250 (1 ~60 600)15 760 447 (1 295000 -4. 8

2499 (8200) 2658 (8720) 6.3

454 458 1.0

1 355 800 (304 800) 1 026 200 (230 700) -24 . 3

1.19 (3.15) 1.57 (4.16 ) 32.1

S change 80% laminar from non-LFC

configuration configura t ion

209 100 (47 000 -13 .0

1 777 065 (399 500 3.3

1 393 628 (313 300 -16.2

58396280312800) -3 . 5

2676 (8780 ) 7.1 I 459 1.1

1 110 700 (249 700 -18.1

1.45 (3 .84 ; 21.9

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(al 100 percent l aminar case.

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Figure 2 . . Drag pola r s with and without laminar f low

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(b) 80 percent lami nar case

Figure 2, ' Concluded,

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Figure 3. - Distribution of airfoil PI·ps.urp corffieipnt

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Figure 4 . .. Ratio 0 airfoil local den ~ i y (I th ,l! 0 frpe ~ rPdl 1l

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Figure 6 . • Di s tribution of a i rfoil local suc tion power coe ff i c i ent.

6

0.430 x 106 1. 92 x 10

. 425 1. 88

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5.90

5.86

5.82

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T . N

(a) 100 pe r cen I an" nar Cd SP .

Figure 7 .. [s imdted ff cts 0 ,a·lpv('l '''' alle,j maxilllulll s tati c th ,-us pe ,- enq ,np on OWL. H1 3W . and range .

1. 34

1. 33

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200 210 220 230 240 250 260 x

(b) 80 percent laminar case.

ri gure 7 , - Co nc ludeo,