Mark Drela MIT Aero & Astroweb.mit.edu/drela/Public/papers/Chicago_10/tasv.pdf · • Global...
Transcript of Mark Drela MIT Aero & Astroweb.mit.edu/drela/Public/papers/Chicago_10/tasv.pdf · • Global...
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