mark drela mit aero & astro

Low-Order Modeling for Conceptual Aircraft Design

and Development of the D8 Transport Concept

Mark Drela

MIT Aero & Astro

Stanford AA295 Seminar

25 May 11

Motivation: NASA’s N+3 Program

N+3 Objective:

Identify concepts and technologies needed

for 70% (!) reduction in Fuel / PAX-mile

from current-technology baseline by 2025

Reasonably safe assumption:

Tweaking current designs will not get us there

Presentation Outline

• Transport Aircraft System OPTimization (TASOPT)

• D8.x “Double Bubble” transport aircraft concept

TASOPT Motivation

Find global optimum in Aircraft + Engine + Ops design space

fuel economy isocontours

existing engines

existing airplanes

common apparent (false) optimum

W/S , M . . .

true optimum

FPR, T

.

.

.

t 4

Pratt’s, GE’s design domain

Boeing’s, Airbus’s design domain

TASOPT Summary - I

Collection of coupled low-order physical models

• Primary structure

– surface beam bending, pressurization loads

• Aero

– viscous/inviscid methods, Trefftz Plane

• Engine

– fan+core 1D flowpath simulation, turbine cooling

• Balance, trim, stability

– surface CL’s at operation and CG limits, static margin

• Performance

– flight trajectory simulation (takeoff, climb, cruise, descent)

TASOPT Summary - II

TASOPT does not use

• Historical correlations for primary structure weights

• Wetted-area× form-factor methods for drag prediction

• Specified-thickness airfoils

• Assumed engine parameters and performance

• Assumed trim conditions

→ These cannot be trusted for “outside the box” designs

Weight Breakdown

fspoi

w.att

flapslat

v.add

h.add

spoiler

aileronl.e., t.e.

wing

fuel

shell

floor

reserve

burn

1

f

cone

1

paddf

1

fstring

frameff fadd

skin

1

reservef

strut

eaddf

1

1

fhadd

1

fvadd

pylon

1

fpylon

window

seat

apu

fff

hpesys

lgnose

lgmain

1

flapffslatff

fribs

lete

f ribs

aile

watt

payload

fuselage

h.tail

v.tail

h.p.e. sysnose l.g.main l.g.

eng.add

eng.bare

nacelleengine

v.capv.web

h.webh.cap

w.add

w.web

w.cap

v.bendh.bend

insulation

web

pay.add

a.p.u.

MTOW

PAYLOAD

STRUCTURE

PROPULSION

EQUIPMENT fixed

CONTROLS

Primary Structure — Fuselage

• pressure vessel with added bending and torsion loads

• bending loads from distributed payload, point tail

p∆

(x)

x

W

Wtail

h

+ Wpay padd+ W +shell

+ hLrMh

(x)v

Lr vMv

N

N ( W )+ W + floorWwindow insul + Wseat

added bending stringers

xwbox

WN fix

Tail Load Bending Moments

• Max tail loads at VNE impart fuselage bending moments

• Moments substantially modified by inertial reaction relief

Lh Lv

wing mass

pitch acceleration yaw acceleration

(x)h

static moment(clamped wing root)

static moment(clamped wing root)

(x)vapproximation

approximation

maxmax

wing mass + yaw inertia

inertial relief

actual moment

Primary Structure — Fuselage

• “double-bubble” pressure vessel

• skin also takes torsion loads from vertical tail

• added longeron area for bending loads

⇒ Pressurization, Loads fully size all structural elements

R fuse

dbw

Lv

v

tτtσ

tσ db

added bending material

tskin

Askin

skincone cone

skin

skin

stringers

fuse

max

A

floor

floor beams

Primary Structure — Wing or Tail Surface

• Multiple linear-taper planform, with or without strut

• Aero loading p(η) via local cℓ factors, center and tip rolloffs

• Weight loading w(η) from structure, fuel, engine

η = 2 y/b

10 o

structural box

ηs

(projected)

Λ

Engine weight alternativeto strut force

∆Lt

∆Lo p(η)

(η)w

NWNWinn out

η

structural cross section

Wing or Tail Surface Cross-Section

• Box beam: bending caps with shear webs

• Non-structural LE/TE fairings, slats, flaps, spoilers

• Box interior defines maximum fuel volume

• Spanwise-constant material stresses

⇒ Bending Moments, Shears, fully size wingbox elements

tcap

twebh

fuelA

wc

box

box

hboxhr

Load Cases used for Sizing Primary Structure

Wing: Net load at Nlift

Tails: Max aero load at never-exceed speed VNE

Tailcone: Max torsion from vertical tail load at VNE

Body skin: Pressurization + Max vertical-tail torsion at VNE

Longerons: Payload loading + Tail loads, at Nland, VNE

Secondary Structure and Equipment Weights

Wfix fixed pilots, cockpit, instrumentation weight

W ′window window weight/length

W ′′insul insulation weight/area

fpadd attendants, seats, galleys, IFE, etc. weight/payload

fgear landing gear weight/MTOW

fhpesys hydraulics, pneumatics, electrical weight/MTOW

fflap flaps, ailerons weight/wingbox-weight

Airfoil Parameterization

Optimum tcstrongly depends on rest of design variables, so . . .

• Family of airfoils over range of tc= 0.09 . . . 0.14

• Each is designed for good Mach drag rise behavior

• Drag polars are precomputed

Gives “rubber airfoil” database:

[cdf , cdp] = F(CL⊥, M⊥ ,

tc , Rec)

Parametric Airfoil Performance Model

Typical Cp(x), M(x, y)

Parametric Airfoil Performance Model

Offline-computed airfoil polars forM⊥ = 0.05 . . . 0.82, cℓ = 0.40 . . . 0.90

Parametric Airfoil Performance Model

Sweep theory with root corrections gives 3D wing profile drag

CDwing= F(CL , M∞ , Λ , t

c, Rec)

Λ V

V

Df

DpDp

shockpotential flow streamline

Cp lc

M fdc

dc p

oc

2ok c

shock

potential flow streamlines

( unswept−shock wing portion )Suns

Induced Drag Model

• Trefftz Plane analysis with wake contraction by fuselage

• Includes tail load from trim calculations

/2by

y’ /2b’y’

y

y’o

yo

Γ y’( )y’

z’

w

Γi−

wn

s’

ii

Fuselage Drag Model

• Potential flow via compressible source line+panel usingA(x)

• BL+wake flow via compressible integral BL method, withlateral divergence via body perimeter b0(x)

• XFOIL-type viscous/inviscid coupling and solution

• Used for fuselage drag and BL Ingestion calculations

CDfuse=

2Θwake

S

x

A(x)∆∗ ∗ΘΘ, , (x)

y

z ∗ ∗, , (x)δ θ θ

b (x)0

Θwake

(x)Λ

(x)λ

Fuselage Drag Model

Typical BL calculation

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

0 10 20 30 40 50 60

HkUe

Uinv

0

0.002

0.004

0.006

0.008

0.01

0.012

0.014

0.016

0.018

0 10 20 30 40 50 60

Disstauw

Knowing BL state at engine location enables BLI accounting

Engine Performance Model

• Complete flowpath simulation for . . .

– . . . sizing at cruise design point,

– . . . operation at off-design takeoff, climb, descent, static.

– Includes effects of Boundary Layer Ingestion (BLI), if any

• Dispenses with engine-performance correlations

• Provides “rubber engine” for optimizer

.m

8

6

0

4

m.

.mα 3

5

7

4.52.5

αc

4.12

1.9

2.11.8

1.8

N NN G/

4.9

l hl

core

core

Turbine Film Cooling Sub-Model

Modified Horlock model, with two prediction modes:

αc = F(Tmetal , Tt 3 , Tt 4 ; StA , θf , ηc) (cooling flow sizing)

Tmetal = F( αc , Tt 3 , Tt 4 ; StA , θf , ηc) (Tmetal prediction)

combustor

T

T

pαc(1− ).m.αc m

.fm

IGV

u

mixingt 5

t 5

Tt 3p t 3

p

T

p

u

p4a

4aTt 4

p t 4

Tt 3u

turbines.m

t f

c

A

B

C

t 4.1

t 4.1

4.1

4.1

TtT TTfawTci Tco m

01

01

0 1η

f

actual adiabatic

y

Tco

T

Tm

Tfaw Tm

Tm

Tci

hypothetical adiabatic

Q.

film edge

θ

θ

g

g

Turbofan Weight Model

Weng =

mcore

100 lbm/s

b0

W0 + W1

OPR

30

b1+ W2

BPR

5

b2

-0.4

-0.3

-0.2

-0.1

0

0.1

0.2

0.3

0.4

0 2000 4000 6000 8000 10000 12000 14000 16000

wei

ght-

pred

ictio

n er

ror

W/W

i - 1

dry weight Wi [lb]

Operation Models — Mission Profiles

• Each design “flown” over specified total range mission to

– . . . size fuel weight,

– . . . evaluate takeoff performance

• Provides “rubber trajectory” for optimizer

γ

cruise−climb descentclimbtakeoff

h

W

WreserveW W W= fuel−dry MTO

CRγDE

totalR

MTOW

W W−MTO burn

0

hc

cR

Operation Models — Takeoff

Analytical model determines normal-takeoff and balanced-fieldlengths ℓTO, ℓBF .

l

V1

V2

l1

V 2

2

2

Aborted Takeofffull power

full power

maximum braking

V 2(l)

V 2(l)

V 2(l)

l

full power one engine outOne Engine−out Takeoff

Normal Takeoff

lBFTO

A

C

B

TASOPT Inputs (Free Parameters)

Wpay payload

Re range

MCR cruise Mach

ℓfield field length

Nlift max flight load factor

Nland max landing load factor

σcap wing sparcap allowable stress

ρcap wing sparcap density

fgear landing gear weight fraction

SMmin static margin at aft-CG case

CL hmintail lift coefficient at forward-CG case

. . . fuselage size and shape

. . . fuselage aero moment function

. . . engine component efficiencies

etc.

TASOPT Operation

Y

DesignVariables

i, j

for i = 1 : Ni for j = 1 : Nj

Configuration, Weight, Fuel burn, T/O perf, ...

Structural gauges

Volumes and Weights

N

Surface spans, areas

Loads, Shears, Moments

Y

Drag, Engine size+weight

Trajectory, Fuel Weight

Total Weight converged?

Optimize Design VariablesN

Fuel burn minimized?

Optimization

Design Closure

Parameter Sampling

DesignParameters i, j

( j )

( i )

( Range, Payload, Mach Nmax, fstress Tmetal, lBFmax ... )

Sweep, CL, AR, Altitude, FPR, BPR, Tt4 ...

( Sweep, CL, AR, Altitude, FPR, BPR, Tt4 ... )

TASOPT Outputs - I

Optimized design variables:

hCR start-of-cruise altitude

CLCRcruise lift coefficient

Λ wing sweep

t/co airfoil thickness at wing root

t/cs airfoil thickness at planform break

λs inner panel taper ratio

λt outer panel taper ratio

rcls cruise cls/clo at break

rclt cruise clt/clo at tip

OPRD design overall pressure ratio

FPRD design fan pressure ratio

BPRD design bypass ratio

Tt 4TO takeoff turbine inlet temperature

Tt 4CRcruise turbine inlet temperature

TASOPT Outputs - II

Airframe:

• Wing and tail areas

• Wingbox location on fuselage

• Structural gauges and weights

• Engine component sizes, areas

• . . .

Operating variables at all mission points:

• Fuel burn

• Speed, Altitude profiles

• Aero variables CL, CD, CLh. . .

• Engine variables n1, n2, FPR, OPR, Tt 3, Tt 5, ufan, ucore . . .

• . . .

TASOPT Calibration/Verification

B737-800 Sizing, Airframe+Ops Optimization

WMTO Wfuel S Λ◦ λt AR CLCRhCR

Actual 171000 39000 1250 25.00 0.250 10.20 0.550 33500Sized only 166001 38474 1229 25.00 0.250 10.20 0.550 33500Optimized 163862 36923 1289 24.33 0.144 10.61 0.530 34070

OptimizedSized only

Effect of Design Space Expansion

Addition of airfoil, engine, trajectory to design spacegives significantly reduced fuel burn

6

6.5

7

7.5

8

8.5

0.7 0.72 0.74 0.76 0.78 0.8

PF

EI

[ k

J / k

g-km

]

Mach

(none)CL, sweep

t/c, AR, CL, sweepfixed opt engine, t/c, AR, CL, sweep

hCR, rubber opt engine, t/c, AR, CL, sweep

Optimum Aircraft Variation vs Mach, Material Stress

Airframe, engine, trajectory all vary on min-fuel airplane

0.72 0.74 0.76 0.78 0.84

5

6

7

8

9

Mach

PF

EI

[K

J/kg

−km

]

fstress

= (blue) 1.0, 1.25, 1.50 (red)

0.72 0.74 0.76 0.78 0.85

10

15

20

25

Mach

AR

L/D

0.72 0.74 0.76 0.78 0.80

10

20

30

Mach

swee

p [d

eg]

0.72 0.74 0.76 0.78 0.80.1

0.12

0.14

0.16

Mach

t/c

0.72 0.74 0.76 0.78 0.80.5

0.6

0.7

0.8

Mach

CL

C

Lper

p

0 50 100

−100

−50

0

0.72 0.74 0.76 0.78 0.83

3.5

4

4.5

5x 10

4

Mach

Fto

[lb

]

0.72 0.74 0.76 0.78 0.8200

250

300

350

400

Mach

u fan

uco

re [

m/s

]

0.72 0.74 0.76 0.78 0.830

35

40

Mach

alt.

[kft]

N+3 Design Problem Statement

Chosen baseline-aircraft mission (Boeing 737-800):

Payload: 38700 lb (215 lb × 180 pax)Range: 3000 nmi

Find:

“Rubber Airframe + Rubber Engine + Rubber Ops”combination which

- has minimum fuel burn over mission

- meets field length, fuel volume constraints

Transport Configuration Considerations

• Weight characteristics

– Fuselage load paths, structure

– Landing gear length

– Wing lift fraction (wing size and weight)

– Stability and CG range (H tail size and weight)

– Engine-out moments (V tail size and weight)

– Nacelle and engine-mount weights

• Drag characteristics

– Stability and CG range (H tail area and drag)

– Engine-out moments (V tail area and drag)

– “Interference” drags

– Nacelle and pylon drag

– Boundary Layer Ingestion?

• Cabin layout

• . . .

Final N+3 Designs

D8.1 (Aluminum) D8.5 (Composite)

B737−8000.80 Mach15.2 L/D166k MTOW8000 ft field

0.72 Mach22.0 L/D130k MTOW5000 ft field

0.74 Mach24.9 L/D100k MTOW5000 ft field

−70% Fuel Burn−49% Fuel Burn

−45% Fuel Burn

D8.1b19.5 L/D120k MTOW

Fuselage Comparison

700 10 20 30 40 50 60 80 90 100 110 120 ft

9

700 10 20 30 40 50 60 80 90 100 110 120 ft

D8.x

B737−800

8

D8.1 Details

70

010

20

30

40

50

60

80

90

100

110

120 ft

22,23 rows180 seats19"x33"

70

010

20

30

40

50

60

80

90

100

110

120 ft

10 20 30 40 50 60

optionalthrustreverser

N+3 D8.1

Dfan = 49 inFPR = 1.58BPR = 7.1OPR = 35.8

8

Mach = 0.72Area = 1298 ft^2Span = 150 ftMAC = 10.6 ftAR = 17.3L/D = 22.0MTOW = 129965 lbWfuel= 20233 lbRange= 3000 nmiField= 5000 ft

Sh=252 ft^2Sv=144 ft^2

D8.x Configuration (vs B737)

Lifting nose, rear flat fuselage

• increased fuselage carryover lift → smaller wing

• built-in nose-up moment from nose lift → smaller tail

D8.x

B737−800

D8.x Configuration (vs B737)

Wide double-bubble fuselage with lifting nose

• partial span loading via 216” wide fuselage (vs 154”)

• reduced floor-beam weight via center floor support

• improved propulsive efficiency via fuselage BL Ingestion

• reduced profile drag via fuselage BL Ingestion

• shorter landing gear and load path

M(y) M(y)

B737 D8.1

D8.x Configuration (vs B737)

Reduced M = 0.72 with unswept wing (vs M = 0.80)

• reduced CDi, via larger AR allowed by unsweep

• LE slat can be eliminated, via increased CLmax from unsweep

• NLF on wing bottom possible, via unsweep and no slat

• faster load/unload of two aisles compensates for slower cruise

D8.x

30 x 6 per aisle

per aisle23 x 4

(30 minutes load,unload)

(15 minutes load,unload)

B737−800

D8.x Engine/Tail Configuration

• Rear fuselage and tails double as flow-aligning nacelles– only minimal nacelles needed

– shield fan faces from ground observers

• Provides Boundary Layer Ingestion (BLI)– local potential flow M ≃ 0.6 matches fan requirement

– no additional BL diffusion – no streamwise vorticity into fan

• Fin strakes synergystically exploited:– function as pylons carrying engine loads and tail surface loads

– shield fan faces from ground observers

D8.1 BL Profiles at Cruise

• local potential flow M ≃ 0.6 matches fan requirement

• little nacelle diffusion – no streamwise vorticity into fan

• only fan does BLI — core intake sees clean flow- avoids core surge risk

- ideal situation for best overall SFC

1 ft

θ

e

k

M = 0.784H = 1.217 = 0.074 ft

θ

e

k

M = 0.724H = 1.265 = 0.108 ft

θ

e

k

M = 0.623H = 1.372 = 0.185 ft

Breguet Parameter Comparison

Wfuel = WZF

exp

TSFC′

M

D′

L

R

a

− 1

≃ WZF

TSFC′

M

D′

L

R

a

737-800 D8.1 D8.1b D8.5

WMTO 166001 129965 120366 99756

Wfuel 38474 20233 21174 11296

WZF 127528 109732 99192 88460

TSFC′/M 0.694 0.628 0.633 0.500

L/D′ 15.18 22.00 19.53 24.85

W WZF fuelM LW WZF fuelM L W WZF fuelM L

1

0

D8.1 D8.5B737−800

TSFC’ TSFC’ TSFC’D’ D’ D’

D8.1b

CD/CL Breakdown by Component

Induced Fuselage WingTOTAL / 2

B737

Tails Nacelles

0.0193

0.01570.0176

0.0085

0.0053

0.00310.0013

0.0664

0.0151 0.0137

0.0184

0.0512D8.1b

D8.1 Fuselage-BLI Benefit Breakdown

No BLI With BLI With BLIηfan = 0.92 ηfan = 0.90 ηfan = 0.92

WMTO 134585 129997 129721

Wfuel 21925 20504 20172

WZF 112660 109493 109549

TSFC′/M 0.6573 0.6367 0.6267

L/D′ 22.03 22.10 22.08

W WZF fuelM LW WZF fuelM L

0TSFC’ TSFC’D’ D’

fanη = 0.92With BLIfan

W WZF fuelM LTSFC’ D’

With BLIfan1.0

0.9

No BLIη = 0.90 η = 0.92

Wing-Body Panel Solutions

737-800 D8.1

Fuselage Centerline Pressures at Cruise

737-800 D8.1

-0.4

-0.2

0

0.2

0.4

0 20 40 60 80 100 120

Cp

x

-0.4

-0.2

0

0.2

0.4

0 20 40 60 80 100 120x

D8.x fuselage has

• larger carryover lift

• positive moment from nose lift

CM Components About Wing+Fuse A.C.

-0.3

-0.25

-0.2

-0.15

-0.1

-0.05

0

0.05

0.1

0.15

0.2

0.4 0.5 0.6 0.7 0.8 0.9 1

CM

_ac

CL

D8.1 fuse

D8.1 wing+fuse

D8.1 wing

B737 wing

B737 fuse

B737 wing+fuse

D8.x fuselage gives ∆CM=+0.125 built-in trim offset benefit

D8.1 Fuselage Lift and Moment Benefits

• Smaller horizontal tail required for forward-CG sizing case

• Smaller trim drag at all flight conditions

• Smaller wing for given required cruise SCL

D8.1 Spanload at Landing, Forward CG

D8.1 Loading at Landing, Forward CG

D8.1 Spanload at Cruise, Mid-range CG

D8.1 Fuselage Lift and Moment Benefit Breakdown

Std. Fuselage D8.x Fuselage∆fLo = −0.3∆CM = −0.125

WMTO 132813 129721

Wfuel 21374 20172

WZF 111439 109549

TSFC′/M 0.6285 0.6267

L/D′ 21.29 22.08

W WZF fuelM LW WZF fuelM L

0TSFC’ TSFC’D’ D’

1.0

0.9

Standard Fuselage D8.x Fuselage

Tail Size and Loads Comparison

D8.1 horizontal tail is 28% smaller and 27% lighter.Two-point support reduces bending moment and weight.

Sh = 350 ft Sh = 252 ft2 2

B737 D8.1

Wh = 2320 lb Wh = 1690 lb

Tail Size and Loads Comparison

D8.1 vertical tail total is 50% smaller and 70% lighter.5x smaller engine-out yaw moment no longer sizes the VT.

Sv = 144 ft 2

Sv = 284 ft 2

B737 D8.1

Wv = 1570 lbWv = 470 lb

Optimum-Engine “Surprises”

• Optimized takeoff turbine inlet temperature is modest→ Excessive takeoff Tt 4 carries cooling-flow penalty in cruise

• Optimized Bypass Ratio and Fan Pressure Ratio are modest→ BLI favors smaller engine and higher fan loading

D8.1 GE90Tt 4TO 1543◦K 1770◦KBPR 7.1 9.0FPR 1.58 1.50

u u2

Fnet

jetE

( ) Fnet( )=

( ) ( )jetE<

A

B

B

B A

A

. .m u.

m u. 21

2

=

=

+

−

+−

+

++

+

B737-800 → “Morphing” Study

Modifications are made sequentially in 8 steps:

0. B737-800, CFM56, M=0.80, lBF ≤ 8000ft, not optimized

1. B737-800, optimized airframe+ops (engine fixed)

2. Fuselage replaced by double-bubble D8.x fuselage

3. Reduced speed to M=0.76

4. Reduced speed to M=0.72

5. Engines moved from wing, to rear fuselage with BLI

6. Optimized airframe+ops+engines, with 15-year engine im-provements

7. Removed slats (less weight, excrescence drag)

8. Reduced field length to lBF ≤ 5000ft

Span constraint for all cases: b ≤ 118 ft

B737-800, optimized airframe+ops, engine fixed

-120

-100

-80

-60

-40

-20

0

20

-100 -80 -60 -40 -20 0 20 40 60 80 100

Case 0Case 1

Fuselage replaced by double-bubble D8.x fuselage

-120

-100

-80

-60

-40

-20

0

20

-100 -80 -60 -40 -20 0 20 40 60 80 100

Case 1Case 2

Reduced speed to M=0.76

-120

-100

-80

-60

-40

-20

0

20

-100 -80 -60 -40 -20 0 20 40 60 80 100

Case 2Case 3

Reduced speed to M=0.72

-120

-100

-80

-60

-40

-20

0

20

-100 -80 -60 -40 -20 0 20 40 60 80 100

Case 3Case 4

Engines moved from wing, to rear fuselage with BLI

-120

-100

-80

-60

-40

-20

0

20

-100 -80 -60 -40 -20 0 20 40 60 80 100

Case 4Case 5

Optimized airframe+ops+engines

-120

-100

-80

-60

-40

-20

0

20

-100 -80 -60 -40 -20 0 20 40 60 80 100

Case 5Case 6

Removed slats for reduced weight and excrescence drag

-120

-100

-80

-60

-40

-20

0

20

-100 -80 -60 -40 -20 0 20 40 60 80 100

Case 6Case 7

Reduced field length to lBF=5000ft

-120

-100

-80

-60

-40

-20

0

20

-100 -80 -60 -40 -20 0 20 40 60 80 100

Case 7Case 8

Total and Fuel Weight Sequence

0

10000

20000

30000

40000

50000

60000

70000

80000

90000

0 1 2 3 4 5 6 7 8

Alt_

toc

[ ft ]

W

eigh

t [ lb

]

Case

737-

800,

M=

0.8

737-

800

opt.

DB

fuse

lage

.

slow

to M

=0.

76

slow

to M

=0.

72

engi

nes

to r

ear

opt.

eng

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Fuel Weight Sensitivity to Span Constraint

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span [ft]

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737 Wfuel737 D / L 737 W_ZFD8.1 WfuelD8.1 D / L D8.1 W_ZF

20:1 Low-Speed Wind Tunnel Model

L/D from Measurement, Re-Correction

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L/D

D8.1 model, 21 Sep 10

80 mph120 mph

tripped 80 mphtripped 120 mph

Measured L/D = 16.1 for tripped model at 120 mphcorrects to L/D = 22.8 at full-scale Re, low MachD8.1 calc. L/D = 22.0 at full-scale Re, M=0.72

Conclusions

• Transport fuel burn minimization in global design space:

– Airframe: t/c, AR, taper, sweep

– Aero: cruise CL, spanload shape

– Engine: fan pressure ratio, bypass ratio, takeoff T4, cruise T4

– Operation: cruise altitude

• N+3 D8.x configuration and optimization gives up to-49% fuel burn with current Al technology + BLI-70% fuel burn with 2035 forecast tech. + BLI

• Physical origins of improvements have been identified

• Preliminary WT data supports predicted performance

Extra Slides

D8.5 Details

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22,23 rows180 seats19"x33"

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optionalthrustreverser

N+3 D8.5Mach = 0.74Area = 1162 ft^2Span = 170 ftMAC = 8.12 ftAR = 24.85L/D = 25.3MTOW = 101590 lbWfuel= 11485 lbRange= 3000 nmiField= 5000 ft

Dfan = 52 inFPR = 1.42BPR = 20OPR = 50

D13.1 (777 size, M=0.72, 1-class, Aluminum)

Dfan = 93 inFPR = 1.55BPR = 8.0

Sh=744 ft^2Sv=590 ft^2

Range= 7600 nmiMach = 0.83Area = 4780 ft^2Span = 213 ftMAC = 26.7 ftAR = 9.5L/D = 21.1MTOW = 640100 lbWfuel= 212500 lb

42,43 rows500 seats19"x33"

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thrustreverser

D13.1

D15.1 (777 size, M=0.83, 1-class, 3-aisle, Aluminum)

Dfan = 93 inFPR = 1.55BPR = 8.0 Sh=709 ft^2

Sv=710 ft^2

22,23 rows500 seats19"x33"

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Range= 7600 nmiMach = 0.83Area = 3650 ft^2Span = 226 ftMAC = 15.8 ftAR = 14.0L/D = 24.0MTOW = 440000 lbWfuel= 98000 lb

D15.1

D17.1 (787 size, M=0.85, 3-class, Aluminum)

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thrustreverser

12 x 14 rows = 168 seats 33"x19"

D17.18 x 6 rows = 48 seats 39"x26"

4,6 x 2 rows = 10 seats 60"x30"

Dfan = 60 inFPR = 1.79BPR = 6.70

Range= 8000 nmiMach = 0.85Area = 2410 ft^2Span = 163 ftMAC = 17.8 ftAR = 11.0L/D = 19.0MTOW = 274000 lbWpay = 50000 lbWfuel= 102500 lb

Sh=344 ft^2Sv=340 ft^2

D17.4 (787 size, M=0.85, 3-class, Composite)

thrustreverser

12 x 14 rows = 168 seats 33"x19"

8 x 6 rows = 48 seats 39"x26"

4,6 x 2 rows = 10 seats 60"x30"

D17.4Range= 8000 nmiMach = 0.85Area = 2240 ft^2Span = 192 ftMAC = 14.0 ftAR = 16.4L/D = 20.8MTOW = 200000 lbWpay = 50000 lbWfuel= 60600 lb

Dfan = 72 inFPR = 1.54BPR = 22.0

Sh=338 ft^2Sv=349 ft^2

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