An Integrated Approach to the Design-Optimization of an N+3 Subsonic Transport

Transcription

1 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

2 Motivation: NASA s N+3 Program Identify concepts and technologies needed for 70% reduction in Fuel / PAX-mile from B baseline by 2025 Tweaking current designs will not get us there

3 Problem Statement B baseline-aircraft mission: Payload: 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

4 Presentation Outline Transport Aircraft System OPTimization (TASOPT) D8.x Double Bubble transport aircraft concept

5 TASOPT Summary - I Collection 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

6 TASOPT does not use TASOPT Summary - II 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

7 Primary Structure Fuselage pressure vessel with added bending and torsion loads bending loads from distributed payload, point tail N ( Wpay + Wpadd + W shell + W window + W insul + W floor + W seat ) added bending stringers r Mv L v p N W tail + r Mh Lh h (x) x wbox v (x) x

8 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 t skin stringers w db added bending material R fuse σ skin t skin L v τ max cone t cone floor σ skin t db v floor beams A skin A fuse

9 Primary Structure Wing or Tail Surface Multiple linear-taper planform, with or without strut Net loading assumed proportional to chord Engine weight alternative to strut force structural box Λ (projected) structural cross section L o p(η) NW inn NW out L t η = 2 y/b 0 ηo η s w(η) 1

10 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 t cap A fuel r h h box h box t web w box c

11 Load Cases used for Sizing Primary Structure Wing: Tails: Net load at N lift Max aero load at never-exceed speed V NE Tailcone: Max torsion from vertical tail load at V NE Body skin: Pressurization + Max vertical-tail torsion at V NE Longerons: Payload loading + Tail loads, at N land,v NE

12 Secondary Structure and Equipment via Fractions W fix f padd f gear f flap pilots, cockpit, instrumentation attendants, seats, furnishings, life support, etc. landing gear flaps, ailerons f hpesys hydraulic, pheumatic, electrical systems

13 Airfoil Parameterization Optimum t c strongly depends on rest of design variables, so... Family of airfoils over range of t c = Each is designed for good Mach drag rise behavior Drag polars are precomputed Gives rubber airfoil database: [c df, c dp ] = F(C L, M, t c, Re c)

14 Parametric Airfoil Performance Model Sweep theory with root corrections gives 3D wing profile drag C Dwing = F(C L, M, Λ, t c, Re c) Λ V V 2 ksuns c o ( unswept shock wing portion ) potential flow streamline D p D p shock c o shock C p cl D f potential flow streamlines M c d f cd p

15 Fuselage Drag Model Potential flow via compressible source line using A(x) BL+wake flow via compressible integral BL method, with lateral divergence via body perimeter b 0 (x) XFOIL-type viscous/inviscid coupling and solution Used for fuselage drag and BL Ingestion calculations b 0 (x) C Dfuse = 2Θ wake S A(x) z δ, θ, θ (x), Θ, Θ (x) y x Θwake

16 Turbofan Performance Model FromKerrebrock,extensivelyenhancedwithvariablec p (T), fan and compressor maps, cooling, etc. π d FPR 0 BPR. m ṁ OPR 4a 2.5 π 3 π b 4 πlc hc 4.5 τht τ 2 α c. m lt Used for... engine sizing engine performance over mission

17 Operation Models Mission Profiles Trajectory equations define mission profiles and fuel burn: W MTO dw dr tanγ = dh dr = F W = F TSFC V cosγ 1 cosγ D L 1 2g d(v 2 ) dr W Wc W e W W = MTO fuel + W reserve W d W reserve W dry = W Wfuel MTO R cruise climb h climb γ c descent takeoff h c h d 0 h b R c R d R h e e R

18 Operation Models Takeoff Analytical model determines normal-takeoff and balanced-field lengths l TO,l BF. full power Normal Takeoff full power One Engine out Takeoff one engine out full power Aborted Takeoff maximum braking V 2 2 V 2 2 V 1 V 2 (l) B V 2 (l) A V 2 (l) C l 1 l TO l BF l

19 TASOPT Inputs (Free Parameters) W pay payload R e range M CR cruise Mach l field field length N lift max flight load factor N lift max landing load factor σ cap wing sparcap allowable stress ρ cap wing sparcap density f gear landing gear weight fraction SM min static margin at aft-cg case C Lhmin tail lift coefficient at forward-cg case... fuselage size and shape... fuselage aero moment function... engine component efficiencies etc.

20 Optimized design variables: TASOPT Outputs - I h CR start-of-cruise altitude C LCR cruise lift coefficient Λ wing sweep t/c o airfoil thickness at wing root t/s s airfoil thickness at planform break λ s inner panel taper ratio λ t outer panel taper ratio r cl s cruise c ls /c lo at break r clt cruise c lt /c lo at tip OPR D design overall pressure ratio FPR D design fan pressure ratio BPR D design bypass ratio T t4to takeoff turbine inlet temperature T t4cr cruise turbine inlet temperature

21 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 C L,C D,C Lh... Engine variables n 1,n 2,FPR,OPR,T t3,t t5,u fan,u core......

22 Final N+3 Designs B D8.1 (Aluminum) D8.5 (Composite) 49% Fuel Burn 70% Fuel Burn

23 optional thrust reverser ft 8 Dfan = 49 in FPR = 1.58 BPR = 7.1 OPR = ,23 rows 180 seats 19"x33" Sh=252 ft^2 Sv=144 ft^2 N+3 D8.1 Mach = 0.72 Area = 1298 ft^2 Span = 150 ft MAC = 10.6 ft AR = 17.3 L/D = 22.0 MTOW = lb Wfuel= lb Range= 3000 nmi Field= 5000 ft ft

24 Breguet Parameter Comparison W fuel = W ZF exp TSFC M D L R a 1 TSFC W ZF M D L R a D8.1 D8.5 W MTO W fuel W ZF TSFC /M L/D B D8.1 D8.5 0 W TSFC D W ZF M L fuel W TSFC D W ZF M L fuel W TSFC D W ZF M L fuel

25 B737 CD/CL Breakdown by Component D TOTAL / 2 Induced Fuselage Wing Tails Nacelles

26 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

27 D8.x Configuration (vs B737) Reduced M = 0.72 with unswept wing (vs M = 0.80) reduced C Di, 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 x 6 per aisle (35 minutes load,unload) D8.x 23 x 4 per aisle (20 minutes load,unload)

28 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

29 Wing-Body Panel Solutions D8.1

30 Fuselage Centerline Pressures at Cruise D Cp x x D8.x fuselage has larger carryover lift positive moment from nose lift

31 0.2 CM Components About Wing+Fuse A.C D8.1 fuse CM_ac B737 fuse D8.1 wing+fuse B737 wing B737 wing+fuse D8.1 wing CL D8.x fuselage has C M = built-in trim offset benefit

32 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 SC L

33 Landing Gear Comparison Shorter gear of D8.x allowed by shorter tail, larger dc L /dα Shorter load path reduces support structure

34 Tail Size and Loads Comparison D8.1 Horizontal tail is 28% smaller and 27% lighter. Two-point support reduces bending moment and weight. B737 D8.1 Sh = 350 ft Wh = 2320 lb 2 2 Sh = 252 ft Sh = 1690 lb

35 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. B737 D8.1 Sv = 284 ft 2 Wv = 1570 lb Sv = 144 ft 2 Sv = 470 lb

36 Optimum-Engine Surprises Optimized takeoff turbine inlet temperature is modest ExcessivetakeoffT t4 carriescooling-flowpenaltyincruise Optimized Bypass Ratio and Fan Pressure Ratio are modest BLI strongly favors smaller engine and higher fan loading D8.1 GE90 T t4to 1445 K 1770 K BPR FPR

37 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 with unswept wing, airfoil, engine, ops changes, give up to 49% fuel burn decrease with conventional technology Physical origins of improvements have been identified