An Integrated Approach to the Design-Optimization

of an N+3 Subsonic Transport

Mark Drela

MIT Aero & Astro

AIAA 28th Applied Aerodynamics Conference

30 Jun 10

Motivation: NASA’s N+3 Program

Identify concepts and technologies needed

for 70% reduction in Fuel / PAX-mile

from B737-800 baseline by 2025

→ Tweaking current designs will not get us there

Problem Statement

B737-800 baseline-aircraft mission:

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

⇒

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

- has minimimum fuel burn over mission

- meets field length, fuel volume constraints

Presentation Outline

• Transport Aircraft System OPTimization (TASOPT)

• D8.x “Double Bubble” transport aircraft concept

TASOPT Summary - ICollection of coupled low-order physical models

• Primary structure

• Aero

• Engine

• Balance, trim, stability

• Flight trajectory

Code operation

• Size wing area, tail volumes, etc.

• Size primary structure elements

• Size engine components, turbine cooling flows

• Trim aircraft at all mission points

• Evaluate mission fuel burn, takeoff performance

• Constrained optimization of design variables to minimize fuel burn

TASOPT Summary - II

TASOPT does not use

• Historical correlations for primary structure weights

• Wetted-area methods for drag prediction

• Fixed wing airfoils

• Assumed trim conditions

• Assumed engine parameters

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

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

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

Rfuse

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

• Net loading assumed proportional to chord

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

• 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 via Fractions

Wfix pilots, cockpit, instrumentation

fpadd attendants, seats, furnishings, life support, etc.

fgear landing gear

fflap flaps, ailerons

fhpesys hydraulic, pheumatic, electrical systems

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

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

Fuselage Drag Model

• Potential flow via compressible source line using A(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

b

A(x)

(x)∆∗ ∗ΘΘ, , (x)

y

z ∗ ∗, , (x)δ θ θ

0

Θwake

Turbofan Performance Model

From Kerrebrock, extensively enhanced with variable cp(T ), fanand compressor maps, cooling, etc.

.m

8

6

0

4

m.

.m

πd

πb3

5

72 4a

πhcπlc ht lt

4.5

2

2.5

αc

4.1

τ τ

FPR

OPRBPR

Used for . . .

• engine sizing

• engine performance over mission

Operation Models — Mission Profiles

Trajectory equations define mission profiles and fuel burn:

dW

dR= −F

TSFC

V cos γ

tan γ =dh

dR=

F

W

1

cos γ−

D

L−

1

2g

d(V 2)

dR

hd

he

hc

hb

cR Rd

cγ

cruise−climb

descent

climb

takeoff

h

W cW

Wd

WreserveW

e

W W= fuel−

MTO

We W W= fuel− W+ reserveMTO

dry MTO

RR

R

0

W

Operation Models — Takeoff

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

l

V1

V2

l 1

V2

2

2

Aborted Takeofffull power

full power

maximum braking

V2(l)

V2(l)

V2(l)

l

full power one engine outOne Engine−out Takeoff

Normal Takeoff

l BFTO

A

C

B

TASOPT Inputs (Free Parameters)

Wpay payload

Re range

MCR cruise Mach

ℓfield field length

Nlift max flight load factor

Nlift 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 Outputs - I

Optimized design variables:

hCR start-of-cruise altitude

CLCRcruise lift coefficient

Λ wing sweep

t/co airfoil thickness at wing root

t/ss 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 . . .

• . . .

Final N+3 Designs

D8.1 (Aluminum) D8.5 (Composite)−49% Fuel Burn −70% Fuel Burn

B737−800

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

Breguet Parameter Comparison

Wfuel = WZF

exp

TSFC′

M

D′

L

R

a

− 1

≃ WZF

TSFC′

M

D′

L

R

a

737-800 D8.1 D8.5

WMTO 166001 129965 99756

Wfuel 38474 20233 11296

WZF 127528 109732 88460

TSFC′/M 0.694 0.628 0.500

L/D′ 15.18 22.00 24.85

W TSFC WZF fuelMDLW TSFC WZF fuelM

DL W TSFC WZF fuelM

DL

1

0

D8.1 D8.5B737−800

CD/CL Breakdown by Component

Induced Fuselage WingTOTAL / 2

B737

D8.1

Tails Nacelles

0.0193

0.01570.0176

0.0085

0.0053

0.01400.0122

0.0146

0.00310.0013

0.0447

0.0664

D8.x Configuration (vs B737)

Wide double-bubble fuselage with lifting nose

• increased optimum carryover lift and effective span, via flat rear fuselage

• built-in nose-up trimming moment, via fuselage lift on nose region

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

• reduced floor-beam weight via center floor support

• improved propulsive efficiency via fuselage BL Ingestion

• shorter landing gear

D8.x Configuration (vs B737)

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

• reduced CDi, via larger AR allowed by unsweep

• need for LE slat eliminated, via increased CLmax from unsweep

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

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

737−800

D8.x

30 x 6 per aisle

per aisle23 x 4

(35 minutes load,unload)

(20 minutes load,unload)

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

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 has ∆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

Landing Gear Comparison

• Shorter gear of D8.x allowed by shorter tail, larger dCL/dα

• Shorter load path reduces support structure

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 Sh = 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 2Sv = 284 ft 2

B737 D8.1

Wv = 1570 lb Sv = 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 strongly favors smaller engine and higher fan loading

D8.1 GE90Tt 4TO 1445◦K 1770◦KBPR 7.1 9.0FPR 1.58 1.50

Conclusions

• Examination of entire Airframe+Engine+Ops design space(TASOPT)

• Global optimization for minimum fuel burn

• N+3 D8.x configuration, reduction in Mach, together withunswept wing, airfoil, engine, ops changes, give up to 49%fuel burn decrease with conventional technology

• Physical origins of improvements have been identified