A CONCEPTUAL DESIGN AND ANALYSIS TOOL FORBLENDED WING BODY AIRCRAFT

Jorrit van Dommelen and Roelof VosDelft University of Technology, 2600GB Delft, The Netherlands

Keywords: Blended Wing Body, Conceptual Design, Design Engine

Abstract

Due to the unconventional nature of the blendedwing body (BWB) no off-the shelf softwarepackage exist for its conceptual design. Thepresent paper details a first step towards the im-plementation of traditional and BWB-specific de-sign and analysis methods into a software tool toenable preliminary sizing of a BWB. The toolis able to generate and analyze different BWBconfigurations. The present paper discusses theresults of three different configurations. Thefirst configuration is an aft-swept BWB with aftmounted engines, the second configuration is anaft-swept BWB with wing mounted engines andthe third configuration is a BWB with forwardswept BWB with wing mounted engines. Theseaircraft comply with the same set of top-level re-quirements and airworthiness requirements.

1 Introduction

Since the introduction of the tube-and-wing(TAW) jetliners in the late 1950s, their range pa-rameter, Mcruise×L/D, has been steadily increas-ing from 13 in 1960 to 16 in the mid 1990s.1

Because designers of TAW aircraft can rely onrealistic solutions from decades of research intothis configuration, it is unlikely that significantgains can be achieved unless radically new tech-nologies are employed. There are various alter-native concepts proposed to the TAW configu-ration. One of those concepts is the blended-wing-body (BWB) configuration, which consis-tently promises to increase the aerodynamic effi-

Fig. 1 Aerodynamic Efficiency - (M0 × L/D)variation with date of entry into service1

ciency by another 25%2, 3 and is the only noncon-ventional concept that airline manufacturers havetaken an interest in (notably Boeing’s 8.5%-scaleBWB demonstrator). Although, on a first glance,the 25% increase might not seem as revolution-ary, it does when one realizes that it took morethan 45 years to achieve this improvement on aconventional TAW jetliner.

For TAW aircraft, there are sufficient airplanedesign handbook methods available to performconceptual design studies. Several off-the-shelfsoftware packages can be used that have thesemethods integrated in a user-friendly computercode (e.g. Advanced Aircraft Analysis (AAA)by DARcorp or Program for Aircraft SynthesisStudies (PASS) by Desktop Aeronautics). How-ever, due to the unconventional nature of theblended wing body there exists no such tool fortheir conceptual design based on a given set of

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JORRIT VAN DOMMELEN AND ROELOF VOS

top level requirements. The present paper detailsa first step towards the implementation of tradi-tional and BWB-specific design methods into asoftware tool that aids the designer in the con-ceptual design of a BWB aircraft.

2 Structure of the Design and Analysis Pro-gram

All calculations are performed from the MAT-LAB environment, using customary written cal-culation modules, centered around a main mod-ule. For a detailed description of each of themodules the reader is referred to Van Domme-len, 2011.4 The present paper summarizes thedesign and analysis modules and discusses ex-amplary the results. The program automaticallydesigns a single BWB aircraft for a given set ofinput parameters and analyzes several key char-acteristics. The program structure is solely basedon a feed-forward structure and is presented inFigure 2.

Initial

sizing

Top level

require

ments

Estimates:

L/D, SFC,

Clmax , etc

Analyzer

S, T, Neng

b, WTO, WOE

Wf

Full aircraft

geometry

Analysis results

Scaled input

vectorMulti Model

Generator

Output

function

Ne

Start

x0, airfoils,

Engine vector,

Fin type

Load aircraft

x0, airfoils,

Engine vector,

Fin type

Fig. 2 Single aircraft calculation structure andblocks

When the program is started, the configura-tion of the aircraft has to be defined by the user.The input consists of four vectors. The first vec-tor describes the shape of the planform and thevertical tail. The second vector is the airfoilvector, which defines the airfoils used at variousspanwise locations as defined in streamwise di-rection. The third vector defines the number ofengines and their location. There are two optionsfor the engine location: body mounted engines

and wing mounted engines. The final vector de-scribes the vertical tail configuration. There arefour options for the vertical tail type. The firstoption is a single body mounted vertical tail. Thesecond option is a twin tail at the trailing edge ofthe centerbody. The third option is the wingletoption, which means that the vertical tail planesare fixed at the wing tips of the aircraft. The finaloption is to use no vertical tail at all.

The first part of the program is the initial siz-ing, indicated in green in Figure 2. Here the wingloading, thrust-to-weight ratio is calculated basedon the method laid out in Roskam, 20055 and datafrom conventional aircraft as well as estimatedblended wing body data based on literature. Af-ter the initial sizing the program constructs disci-pline specific models of the aircraft in the multi-model generator module of the program. Themodels are subsequently used to analyze the air-craft on various disciplines, see Section 2.3. Theoutput module post-processes the data and gener-ates a variety of plots. Each of these modules aredescribed in more detail in the next sections.

2.1 Preliminary Sizing Module

Preliminary sizing of thrust and wing loadingas well as a Class I weight estimation are car-ried out according to traditional handbook meth-ods.6 Estimates of BWB aircraft from previ-ous studies2, 3, 7 are taken into account along withconventional aircraft data where necessary. Fora given set of range requirements and assump-tions on aerodynamic and engine performance,this module calculates the take-off weight, fuelweight, and operational empty weight of the air-plane. With appropriate choices, based on priorresearch into BWB aircraft, for the design max-imum lift coefficients and wing aspect ratio, thedesign point in the the thrust loading vs. wingloading diagram is calculated. The design pointdetermines the take-off thrust and the wing area.Through the aspect ratio, the wing span is alsofixed. The input vector describing the planformof the aircraft is scaled to match the wing areaand wing span estimations of the initial sizing.This approach is used to increase the probability

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A Conceptual Design and Analysis Tool for Blended Wing Body Aircraft

to arrive at a feasible aircraft design at the end ofthe conceptual design phase.

2.2 Multi Model Generator Module

The multi model generator (MMG) uses the out-put from the preliminary sizing along with speci-fied user input on the wing geometry, the selectedairfoils, the engine configuration and the verticaltail configuration to generate the following mod-els:

• geometric model of the outer shell includ-ing disposition of vertical tail and engines

• model of the interior volume: passengercabin, cargo space and fuel tank

• structural disposition of the wing box

• vortex lattice model of the wing planform

Figure 3 displays all data generated by the modelgenerator. This data is used in the analysis mod-ules of the program.

2.2.1 Geometric Model of the Outer Shell

The geometric model of the outer shell consistsof three parts: the aerodynamic surface of thewing-body combination, the vertical tail surfacesand the engines. The multi model generator startsby generating the aerodynamic surface of thewing-body combination. This surface is formedby reading the scaled input vector. The wing-body surface is divided in several trunks. Theshape is described by the root chord and five ad-ditional parameters per wing trunk, displayed inTable 1.

Input Variable SymbolWing input Root chord c1Input per trunk Trunk’s tip chord cn+1

Trunk’s tip span bn+1Twist at cn+1 εnSweep trunk n ΛnDihedral trunk n Γn

Table 1 Wing-body input parameters

The planform is formed by defining the 2Dplanform and subsequently adding dihedral andtwist. The aerodynamic surface is formed by alofting process. The airfoil at each trunk end isdefined in the input vector. By interpolating theairfoils, the aerodynamic surface of the BWB isformed.

The definition of the vertical tail is similar tothe wing-body definition. Two additional inputparameters are needed to fix the location of thevertical tail. These are the y-position with respectto the centerline and the distance from the lead-ing edge. The tail is automatically fixed to bodyin z-direction by the multi-model generator. Gen-erally speaking, only a single trunk is needed todescribe the vertical tail. Not all vertical tail in-put parameters are used for certain tail configu-rations. An example is a winglet configuration.In that case, the tail root chord is set equal to thewing tip chord.

The airfoils used for the wing-body are basedon the whitcomb supercritical airfoil. The cam-ber is removed, while the thickness distribution ispreserved. The thickness is controlled by the in-put vector. The vertical tail airfoils are symmetricNACA 4-series profiles with a thickness of 14%at the root and 12% at the tip.

The final part of the outer shell is formed bythe engines. Two options for the engine positionsare considered. The engines can be placed at theaft-body or beneath the wings. The position andnumber of engines is described by the engine in-put vector. The multi-model generator finds theexact location of the engine by considering theaerodynamic surface of the wing-body and the lo-cation of the vertical tail.

2.2.2 Cabin and Fuel Volume

The internal model consists of the fuel tanks andthe pressure cabin. The fuel tanks are located inthe outer trunks of the wing. In this study thefuel tanks are assigned to the two outboard wingtrunks. The fuel tanks are located inside the tor-sion box and extend to 85% of the semi wingspan.

The pressure cabin is formed by a separate

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JORRIT VAN DOMMELEN AND ROELOF VOS

structure, which has to withstand pressurizationloads. Liebeck2 proposed two concepts for thepressurized cabin for BWB. One with an inte-grated skin and pressure shell. This requires athick sandwich structure for both the upper andlower wing surfaces, which takes both aerody-namic loads and pressurization loads. A secondapproach is a separate pressurization structure,which carries only pressurization loads. Thispressure vessel is thin-walled and is loaded solelyin tension. The latter approach is used in thisstudy. This structures takes the form of a multi-bubble structure. By using a separate structure,the functions of the aerodynamic structure andthe cabin are separated. the multi-model gener-ator determines the geometry of the multi bubblepressure cabin. The structure consists of straightcylinders, which are joined together.

2.2.3 Structural Model of the Wing-Body

The third model created by the multi model gen-erator is a structural model of the wing-bodycombination. This model forms the basic shapeof the wingbox, which carries the loads throughthe structure. The front and rear spar are assumedto be located at fixed chord percentages at 13%and 72%, respectively. This model is used forthe Class II weight estimation. The weight ofthe wing-body structure is determined by consid-ering the wingbox as the load carrying structurewhich is statically determined.

2.2.4 Vortex Lattice Model

The multi model generator generates a pro-gram specific model for the aerodynamic analy-sis module. A vortex lattice method is employed.The aircraft consist of the wing-body and the ver-tical tail surfaces and is modeled with zero thick-ness. This model is used for the calculation ofthe lift and drag polar in cruise flight along withstability derivatives.

2.3 Analysis Module

After the full geometry of the aircraft is gener-ated, a set of modules analyzes key characteris-tics of the airplane. The first module is the pay-

load module, which uses the cabin layout to ana-lyze the volume and area available for the pas-senger cabin and the cargo bays. The secondmodule performs a basic aerodynamic analysisbased on a vortex lattice method. When the aero-dynamic loads are known, a detailed (Class II)weight estimation is performed and the center ofgravity (CG) travel during loading and flight isdetermined. The landing gear position and strutlengths are calculated based on the CG range,clearance and loading requirements. Combiningthe CG and aerodynamic information, the trimdeflection in cruise flight is estimated along withestimates for the trim drag. In addition, estimatesare made for the static longitudinal, directionaland lateral stability, take-off rotation capabilityand OEI controllability. Subsequently, the per-formance module calculates the payload-rangediagram and estimates the take-off field length,the landing length, and the attainable climb gradi-ents for various engine operating/non-operatingconditions. The final module compares the out-put from each of the analysis modules to the top-level requirements and aviation regulations. Thisanalysis block is visualized in Figure 4 and dis-cussed in detail below.

Aerodynamic

analysis

Full aircraft

geometry

Analysis results

Cabin layout

analysis

Class II

weight

estimation

Empty

aircraft CoG

Fuel and

payload CoG

Trim and

stabilityPerformance

Constraints

analysis

Fig. 4 Analysis block flow diagram

2.3.1 Payload Module

The payload module analyzes the pressure cabingenerated by the multi model generator. The re-quired amount of passenger cabin floor area andthe required cargo volume are determined fromthe top-level requirements. The payload mod-ule determines the location of the passenger cabinfloor, see Figure 5.

Subsequently, the payload module calculatesthe available floor area and cargo volume. The

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A Conceptual Design and Analysis Tool for Blended Wing Body Aircraft

Aircraft

Geometry

Outer Shell

Model

wing-body

surface

wing

trunk 1

wing

trunk 2

wing

section 1

wing

section 2

wing

section 1

...

....

tail assemblytail

tail plane 1

...

root section

tip section

engines

engine 1

engine 2

...

Interior Model

fuel tanksel tanks

fuel tank

trunk 1

...

cabin

cylindrical

section

l

bulkheads

floors

Structural

Model

wing

structure

wingbox

trunk 1

front spar

rear spar

wingbox

trunk 2

Votex Lattice

Model

wing-body

wing

trunk 1

wing

trunk 2

wing

section 1

wing

section 2

wing

section 1

...

....

tail assembly

tail plane 1

...

root section

tip section

Fig. 3 Graphical overview of data generated by the MMG

Pressure shell and !oors

Centerline

Aerodynamic cross−section

Bubble centers

20

0 c

m1

63

cm

10

cm

Passenger compartment

Cargo compartment

Fig. 5 Cross-section of the cabin, showing pas-senger and cargo compartments

vertical dimension of 1.63m is considered to cal-culate the available cargo volume, to give the pos-sibility to accommodate the most common LDtype cargo containers and pallets.

2.3.2 Aerodynamic Module

The second analysis module is the aerodynamicmodule. The aerodynamic module uses the vor-tex lattice method to determine the lift polar forcruise conditions. The determination of the in-duced drag is done using a Trefftz plane analysis.

For the determination of the zero-lift drag, themethod of Raymer is used.8 This method usesthe flat-plate friction drag and form factors to es-timate the zero-lift drag.

Secondly, the aerodynamic module estimatesthe maximum lift coefficient. First the maxi-mum lift coefficient in clean configuration is de-termined. This is done by taking the sections’maximum lift coefficient into account. Thesevalues are determined from experimental data.9

The spanwise lift distribution is interpolated untilthe sections’ maximum lift coefficient is reached.The corresponding lift coefficient is consideredto be the maximum lift coefficient of the aircraft.The effect of the addition of slats, which are as-sumed to be the only high-lift devices present, isdetermined from handbook methods.

The final part of the aerodynamic module isto estimate the critical Mach number and theDrag Divergence Mach number. These values areused to conclude if the design is considered fea-sible. The relation between quarter chord sweep,

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JORRIT VAN DOMMELEN AND ROELOF VOS

local thickness and lift coefficient from Kroo10 isused to estimate the critical Mach number. Thedrag divergence Mach number is estimated byadding 0.05 to the critical Mach number.

2.3.3 Weight Module

With the aerodynamic parameters estimated, asecond, more detailed weight estimation is per-formed. This weight estimation is based on theactual dimensions of the aircraft, whereas theClass I weight estimation was solely based onstatistical and empirical relations.

The take-off weight is kept constant to thetake-off weight calculated in the initial sizing, asis the payload weight. This means that emptyweight is traded for fuel weight. This approachmeans that no iterations in design weight are re-quired.

Traditionally, empirical relations between theaircraft dimensions and components weight areused in Class II weight estimation. This is suf-ficiently accurate for conventional aircraft, butcannot be applied to the structural componentsof unconventional aircraft. The primary struc-tural component is the wing-body structure. Theweight is estimated using the method of Toren-beek.11 This method assumes that all loads willbe ultimately concentrated in the primary wing-box structure, consisting of upper and lower stiff-ened panels, a front and rear spar and ribs. Themethod assumes that the wingbox is staticallydetermined, where bending forces are absorbedby the stiffened skin panels and shear forces aretransferred by the spar webs. Torsional loads arenot taken explicitly into account. Maneuveringand gust loads are considered as the determin-ing load cases. Correction factors are applied tothe idealized structure to account for non-taperedskins, splices, joints and access panels. The wingbox weight is determined per wing trunk, usingthe structural model of the wing-body as gener-ated by the multi-model generator. The bendingmoment and shear forces are calculated from thelift distribution, which is determined in the aero-dynamic module. The weight of the secondarywing-body structure, the empirical relations from

Torenbeek11 are used.The pressure cabin forms a separate struc-

ture, which is assumed to absorb the load causedby the pressure difference only. When using amulti-bubble structure as described by Geuskenset al.,12 the structural weight is independent ofthe shape. A Correction factor of 2 is appliedto account for the absence of continuous verti-cal walls between the bubbles of the structure. Asecond correction is applied for the existence ofdoors in the structure according to Raymer.8

The weight of the vertical tail and the weightof the aircraft’s systems is determined using themethod of Raymer.8 The use of this method ispossible since the systems do not differ funda-mentally from conventional aircraft.

2.3.4 Center of Gravity Module

With the aircraft component’s weight known, theprogram determines the center of gravity of theempty aircraft. The geometrical model is used toestimate the position of the components.

The center of gravity at each point in flightdepends on the amount of payload, the amount offuel and the way the fuel is loaded. These load-ing conditions are taken into account by dividingthe payload-range diagram in 31 points. Each ofthese points on the horizontal axis is consideredto be a missions. Each mission is divided in 109phases. For the flight phases other than the cruiseflight, fixed fuel fractions are used. The remain-ing fuel fraction for the cruise phase is dividedin 100 phases. A total of 4 different fuel loadingschemes are considered. The fuel tanks are lo-cated in the two outboard wing trunks. Loadingcombinations are formed by filling either the in-board or the outboard tank completely, and startusing the fuel from either one of these fuel tanksfirst. The center of gravity is calculated for eachof these situations and is stored in a 31x109x4matrix.

Based on the center of gravity range occur-ring in ground operations, the landing gear po-sition is determined. The rules as described byRaymer8 are used. Based on the clearance mar-gins of the aft end of the body, the engines and

6

A Conceptual Design and Analysis Tool for Blended Wing Body Aircraft

the wing tips, the length of the landing gear isdetermined.

2.3.5 Trim and Stability Module

The trim and stability module starts by determin-ing the trim position for each center of gravitycase. To be trimmed in longitudinal direction,the lift must equal the drag and the aerodynamicmoment must equal the moment generated by theaircraft’s weight. This creates a system of twoequations, which can be solved to determine thetrim deflection of the pitch control surfaces andthe angle of attack. The aerodynamic derivativesas determined by the aerodynamic module areused. Using the angle of attack and the trim de-flection, the trim drag is estimated. The trim dragis found by considering the derivative of CDδ withrespect to α. Multiplying this derivative by thetrim deflection δ and the angle of attack α, givesan indication for the increase in drag coefficientdue to the elevator deflection.

The static longitudinal stability is determinedby calculating the minimum static margin. Thestatic margin is the difference between the cen-ter of gravity position and the neutral point. Theneutral point is found by dividing the differencein aerodynamic moment for two different anglesof attack, by the difference in lift coefficient atthese angles of attack. These moments are al-ready known from the aerodynamic module. Theminimum static margin is determined by consid-ering the distance between the most aft center ofgravity position and the neutral point, expressedin terms of the mean aerodynamic chord (MAC).In the present study some instability is allowed:the minimum static margin is set at -10%.

The directional and lateral stability are takeninto account by considering the weathercock sta-bility and the rolling moment due to sideslip, re-spectively. The coefficients are obtained from theaerodynamic module, with a correction for theabsence of thickness in the vortex lattice model.Additionally, the stability module calculates theone-engine-inoperative trim under the most un-favorable circumstances, using the method inRoskam.9 The take-off rotation speed with the

pitch control surfaces in take-off rotation deflec-tion is calculated by considering the equilibriumsituation.

2.3.6 Performance Module

With all necessary data known, the performancemodule refines the performance estimates. Thecruise performance is evaluated by calculatingthe lift-to-drag ratio using aerodynamic data andtrim drag from the trim and stability module. Therange for each cruise flight phase is calculated,to construct the payload-range diagram. Thetake-off and landing distance are calculated us-ing Raymer’s balanced field length method.8 Theclimb performance is calculated from the excesspower at maximum continuous engine thrust atsea-level.

2.3.7 Constraints Module

The constraints module compares the calculatedcharacteristics of the aircraft with top-level re-quirements and aviation regulations. The con-straints are presented in Table 2.

Table 2 Constraints summaryParameter Constraints UnitWing span b < 80 mAircraft overall length l < 80 mCabin floor area Scabin ≥ Scabinreq m2

Cargo volume Vcargo ≥Vcargoreq m2

Take-off distance sTO ≤ sTOreq mLanding distance sland ≤ slandreq mStall speed take-off VstallTO ≤Vstall,TOreq m/sStall speed clean Vstall ≤Vstallreq m/sClimb gradient OEI1 γOEI1 > 0.012 −Climb gradient OEI2a γOEI2a > 0.000 −Climb gradient OEI2b γOEI2b > 0.000 −Climb gradient OEI2c γOEI2c > 0.012 −Climb gradient OEI3 γOEI3 > 0.032 −Climb gradient AOE1 γAEO1 > 0.021 −Maximum trim deflection δh < 12 degMinimum static margin SM >−10 %Directional stability Cnβ = 0.010 −Dihedral effect Clβ < 0 −Take-off rotation speed Vrot <VstallTO m/sDrag divergence Mach MDD ≤ Mcruise −Nose wheel load 0.05 < Fnlg < 0.20 −Nose landing gear position xnlg > xnose +0.5 mMain landing gear position xle < xmlg < xte m

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JORRIT VAN DOMMELEN AND ROELOF VOS

3 Results and Discussion

By operating the MMG and subsequent analysismodules, the designer can almost instantly ana-lyze a certain blended wing body configurationand compare it to the set of constraints of Table2. This allows the designer to generate a numberof different BWB configurations based on a set of31 input parameters and a set of predefined air-foils and compare their respective performancein order to make a judgement on which designis most promising. Examples of various BWBconfigurations are presented in Figure 6. In ad-dition, the input of the geometry module and theoutput of the analysis module is structured suchthat for a chosen configuration an optimizer canbe employed to find a feasible (or even optimal)BWB geometry that fulfills all the top-level re-quirements and FAR/CS constraints. In the fol-lowing subsections three examples are presentedthat are exemplary for the capability of the designtool.

The aircraft is designed to carry an equal pay-load of 400 passengers and 22,000 kg of freight.The maximum take-off weight of the three air-craft is equal, which allows an objective compari-son based on cruise range, empty weight and con-straints. The cruise speed is Mach 0.82 and thelanding and take-off runway length should be lessthan 2,500 m. In the subsequent sections, threedifferent BWB aircraft are designed for these re-quirements and sequently compared.

3.1 Example: Design of a ‘Conventional’BWB

The most conventional BWB configuration fromprevious studies has aft-swept wings, two verticaltails at the wing tip (doubling as winglets), andthe engines positioned close to the trailing edgeof the center section (see Figure 10). When tryingto find a solution for this configuration that satis-fies all the constraints of Table 2, various con-straints appear to be prohibitive. The most im-portant aspect of this configuration are discussedin this section. The input parameters used to con-struct this concept are given in Figure 7.

Input vector

Sect 1 Sect 2 Sect 3 Sect 4 Sect 5 Sect 6

Chords 39.7 39.1 35.1 23.2 12 5.26 m

Spans 1.18 3.58 8.86 14.7 32.4 m

Twists 1.54 −0.777 °

Sweeps 17.6 57.6 48 48 48 °

Dihedral 0.606 −0.545 1.03 2.01 4.61 °

t/c 0.188 0.176 0.157 0.157 0.142 0.142 %

Tail h 8.67 m

Fig. 7 Conventional BWB configuration, aftmounted engines input

A summary of the most important perfor-mance and constraint parameters is given in Ta-ble 3 in Section 3.4. The wing area and wingloading of all three aircraft lie close to each other.The wing area of the ‘conventional’ BWB is 1091m2, which gives a wing loading of 3650 m/N2.The empty weight of the aircraft is 190.4 103 kg,which is the highest of the three aircraft. The rea-son for this is the lift distribution, which gives ahigher bending moment and causes a larger struc-tural weight of the wing-body structure. Sincethe payload weight and the maximum take-off ofall aircraft are equal, this leaves a smaller amountof fuel weight compared to the other concepts.

This concept has the highest lift-to-drag ra-tio. The maximum lift-to-drag ratio is 27.9, withan average lift-to-drag of 27.2 for a maximum-payload mission. This concept is aerodynami-cally very efficient. The cause is the favorable liftdistribution and the presence of the winglets. Thefuel consumption is 0.0207 kg/pax/km, which is10% lower compared to the other concepts. Themaximum-payload range is 14,359 km and theferry range is 17,670 km. The main influences onthe maximum-payload range are the fuel weightand the aerodynamic efficiency.

The ‘conventional’ BWB has the lowestmaximum lift coefficients and consequently thelongest take-off and landing distances. The liftdistribution is unfavorable for low speed flight.Tip stall is likely to occur and the lift distributionis unfavorable for the maximum lift coefficient.The aircraft has a minimum static margin of -8.43%, which occurs when flying without pay-load and at maximum fuel weight. With these

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A Conceptual Design and Analysis Tool for Blended Wing Body Aircraft

Fig. 6 Example of Various BWB Configurations Generated by the MMG

10

20

30

40

−30

−20

−10

0

10

20

30

0

5

10

10 20 30 40

0

5

10

−30−20−100102030

0

5

10

10 20 30 40

−30

−20

−10

0

10

20

30

Fig. 8 Conventional BWB configuration, aftmounted engines

conditions the center of gravity is shifted far aft,causing the aircraft to be unstable. The aircraftturns out to be instable during most parts of theflight envelope. The shift in center of gravity ex-pressed in terms of the MAC is the largest of thethree concepts. This is mainly caused by the po-sition of the fuel tanks relative to the payload andempty weight’s center of gravity.

3.2 Example: Design of a BWB with wing-mounted engines

The second concept is very similar to the firstconcept. The input parameters used to constructthis concept are given in Figure 9. The engineposition is changed from aft mounted engines, towing mounted engines. The reason for this is toinvestigate the effect of the engine position on thelongitudinal stability of the aircraft. The previ-

Input vector

Sect 1 Sect 2 Sect 3 Sect 4 Sect 5 Sect 6

Chords 40.1 39.4 34.7 23.2 13.4 3.62 m

Spans 1.01 3.3 8.4 13.3 32.1 m

Twists 3 2.53 °

Sweeps 28.6 64.7 47.7 46.2 46.2 °

Dihedral 0.348 −0.851 0.929 0.744 5 °

t/c 0.178 0.177 0.174 0.13 0.13 0.13 %

Tail h 8.03 m

Fig. 9 Conventional BWB configuration, wingmounted engines input

ous concept showed a minimum static margin of-8.43%, while this concept has a minimum staticmargin of only -2.47%. Moreover, the conceptturns out to have a positive static margin through-out most of the flight envelope. For a maximum-payload range mission, the aircraft is always sta-ble. The only instability occurs during a ferryrange mission, at the very start of the cruise flight,with almost full fuel tanks.

Compared to the previous concept, this con-cept has a lower lift-to-drag ratio. The maximumlift-to-drag ratio is 25.5 and the average lift-to-drag ratio in cruise flight is 23.7. The lower aero-dynamic efficiency is caused by the lift distribu-tion, leading to a 10% higher fuel consumption.The lift distribution tends to unload the wing tips.This, in turn, is favorable for the structural weightand low speed performance. The empty weight is182.5 103 kg, which is smaller compared to theprevious concept. The impact of the lower aero-dynamic efficiency on the range is limited by thesmaller empty weight. The smaller empty weight

9

JORRIT VAN DOMMELEN AND ROELOF VOS

5

10

15

20

25

30

35

40

−30

−20

−10

0

10

20

30

−2

0

2

4

6

8

10 20 30 40

−4

−2

0

2

4

6

8

−30−20−100102030

−4

−2

0

2

4

6

8

10 20 30 40

−30

−20

−10

0

10

20

30

Fig. 10 Conventional BWB configuration, wingmounted engines

means that more fuel is available. The maximum-payload range is 13,599 km and the ferry range is14,273 km. The latter is relatively small, becauseof the limited size of the fuel tanks. The maxi-mum lift coefficients in clean and slats extendedconfiguration are 1.47 and 1.63 which is con-siderably higher than the previous concept. Thetake-off and landing distances are smaller and are1,673 m and 2,230 m, respectively.

The most notable aspect of this concepts arethe nose load. The variation is very large andthe nose loads are at the limits of 5% and 20%of the weight respectively. This is caused by theshort-coupled nature of the Blended Wing Body.The nose landing gear is close to the nose, andcannot be moved forward to increase the marginsof the nose loads. The rudder trim deflection inone-engine-out condition is much larger than theother two concepts, which due to the fact that thecritical engine is located further outboard.

3.3 Example: Design of a Forward SweptBWB

A more unusual BWB configuration that has beenanalyzed is a forward-swept concept with en-gines hanging from pylons under the wing (seeFigure 12). The input parameters used to de-fine this concept are given in Figure 11. Theforward sweep can exhibit considerable benefitsin transonic flow, due to the increased shock-

Input vector

Sect 1 Sect 2 Sect 3 Sect 4 Sect 5 Sect 6

Chords 36.9 36.1 22.1 19.7 15 2.58 m

Spans 1.07 8.59 11.1 16.7 31.4 m

Twists −0.228 4.02 °

Sweeps 29.5 64.6 −28.6 −28.6 −28.4 °

Dihedral 0.88 3.97 0.505 1.7 5 °

t/c 0.189 0.189 0.18 0.18 0.15 0.15 %

Tail h 8.34 m

Fig. 11 Forward-Swept BWB configuration input

wave sweep, the reduction of tip-stall tendencies,and successful implementation of natural laminarflow technology.13 Even though none of these ef-fects can be quantified in the conceptual designstage, it is of interest to see if such a configura-tion could meet the constraints of Table 2. Theengines are located beneath the wing for longi-tudinal stability reasons. When using the sameplanform with the engines mounted on the aft-body, the aircraft turned out to have a minimumstatic margin of -47%, effectively ruling out thisconfiguration.

From the planform the decreased leadingedge sweep over the previous concept can be ob-served. This confirms the expectations that lessleading edge sweep is required for forward-sweptwing aircraft compared to aft-swept wing air-craft, to meet divergence Mach number require-ments. The other most notable aspect is the verylow operational empty weight. The operationalempty weight is only 156.7·103 kg. Since themaximum take-off weight of the three conceptis kept equal, the fuel weight is relatively large.This, in turn, means a very large range. Themaximum-payload range is 16,452 km and theferry range is 21,875 km. The aerodynamic effi-ciency is somewhat lower than the aft-swept wingconcepts with the maximum lift-to-drag ratio be-ing 23.7 and the average lift to drag ratio in cruiseflight being 22.5. The fuel consumption is al-most equal to the previous concept and is 0.0235kg/pax/km.

The empty weight of the aircraft is small be-cause of the lift distribution. The lift-distributionis is also favorable at low speed conditions. Stallis likely to occur on the inboard parts of the

10

A Conceptual Design and Analysis Tool for Blended Wing Body Aircraft

5

10

15

20

25

30

35

40

−30

−20

−10

0

10

20

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−2

0

2

4

6

8

5 10 15 20 25 30 35 40

−5

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5

−30−20−100102030

−5

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5

5 10 15 20 25 30 35 40

−30

−20

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10

20

30

Fig. 12 Forward-Swept BWB configuration

wings and the maximum lift coefficient is rel-atively high. The maximum lift coefficients inclean and slats extended configuration are 1.44and 1.58 respectively, which is close to the valuesof the previously discussed concept. The take-offand landing distances are similar as well.

The minimum static margin lies in betweenthe previous two concepts and has a value of -5.54%. Since the fuel’s center of gravity, the pay-load’s center of gravity and the empty weight’scenter of gravity are located close to each otherthe center of gravity travel is small compared tothe aft-swept wing concepts. The center of grav-ity travel is only 9.41% of the MAC. This alsomeans that the trim variation is less and thereforethe maximum trim deflection is less than one de-gree. The size of the control surface may be re-duced for this type of configuration.

3.4 Concept Comparison

In Table 3 the results of the optimized aircraft aredisplayed. Values printed in blue highlight thebest values, while values printed in red highlightthe worst values.

4 Conclusion and Future Work

A conceptual design tool for the configuration de-sign of a blended wing body (BWB) has beenpresented. The capability of this tool allows thedesigner to quickly alter the main geometric com-ponents of the airplane and analyze its effect onkey characteristics such as flight and field perfor-

Parameter Symbol UnitWing area S 1091 1043 1074 m2

Wing span b 64.9 64.4 62.7 mAspect ratio A 3.86 3.98 3.67 -Wing loading MTOW W/S 3650 3818 3706 N/m2

Overall length - 41.4 40.5 38.5 mMTOW WTO 405.9 405.9 405.8 103 kgOEW WOE 190.4 182.5 156.7 103 kgPayload weight Wpl 66.4 66.4 66.4 103 kgFuel weight Wf 149.1 157.0 182.7 103 kgMax-payload range Rmaxpl 14359 13599 16452 kmFerry range Rferry 17670 14273 21875 kmFuel consumption - 0.0207 0.0234 0.0235 kg/pax/kmMax lift-to-drag ratio (L/D)max 27.9 25.5 23.7 -Av lift-to-drag ratio (L/D)av 27.2 23.7 22.5 -Max lift coefficient CLmax,clean 1.12 1.47 1.44 -Max lift coeff, slats CLmax,land 1.28 1.63 1.58 -CG travel - 14.9 13.6 9.41 % MACCabin floor area Scabin 312 276 273 m2

Cargo volume Vcargo 186 216 239 m3

Take-off distance sTO 2006 1673 1655 mLanding distance sland 2477 2230 2249 mStall speed TO land VstallTO 62.6 56.6 56.7 m/sStall speed clean Vstall 66.7 59.7 59.3 m/sClimb gradient OEI1 γOEI1 7.4 5.5 5.0 %Climb gradient OEI2a γOEI2a 4.0 2.0 1.4 %Climb gradient OEI2b γOEI2b 7.4 5.5 5.0 %Climb gradient OEI2c γOEI2c 10.6 8.7 8.3 %Climb gradient OEI3 γOEI3 17.0 16.7 16.4 %Climb gradient AOE1 γAEO1 16.6 15.6 15.1 %Max trim deflection δe,max 4.5 4.8 0.9 degMin static margin SM -8.43 -2.47 -5.54 % MACWeathercock stability Cnβ 0.0524 0.0311 0.0747 1/radEffective dihedral Clβ -0.138 -0.100 -0.0329 1/radTake-off rotation speed Vrot 30.8 31.1 27.1 m/sOEI rudder deflection δr,OEI 6.24 21.2 8.41 degDrag div Mach out tr MDD 0.93 0.90 0.84 -Drag div Mach out tr-1 MDD 0.82 0.82 0.82 -Nose landing gear x xnlg 1.99 1.90 1.99 mMain landing gear x xmlg 26.9 26.0 19.2 mMinimum nose load - 6.3 5.0 8.5 %Maximum nose load - 18.7 20.0 16.5 %

Table 3 Concept comparison table

mance, static stability, balance, weight, and pas-senger accommodation. The tool includes meth-ods for the sizing of all components (body, wing,fin, landing gear) relevant to this design stage andrelies on modified handbook methods for BWB-specific components, such as the structural lay-out of the cabin, the structural weight estimationand aerodynamic analysis of the aircraft.

Future work includes the refinement of vari-ous analysis methods, tailored to increase the fi-delity of the analysis modules. The present workis to be integrated in a conceptual design andanalysis tool that allows the designer to gener-ate conventional and nontraditional airplane con-figurations such as a box-wing configuration anda blended-wing-body configuration. Finally, thistool is to become the first part in a rule-based de-

11

JORRIT VAN DOMMELEN AND ROELOF VOS

sign framework in which the fidelity of the ge-ometric and analysis models increases progres-sively with each design step.14 The robust imple-mentation of the present work within this frame-work is the subject of future investigations.

Acknowledgements

The authors would like to thank Dr. E. Torenbeekfor sharing his insights on non-conventional air-plane configurations and introduction to relevantliterature.

References

[1] Denning, R., Allen, J., and Armstrong, F., “Thebroad delta airliner,” The Aeronautical Journal,Vol. 107, No. 1075, September 2003, pp. 547–558.

[2] Liebeck, R., “Design of the Blended WingBody Subsonic Transport,” Journal of aircraft,Vol. 41, 2004, pp. 10 – 25.

[3] Hileman, J. I., Spakovszky, Z. S., Drela, M.,Sargeant, M. A., and Jones, A., “Airframe De-sign for Silent Fuel-Efficient Aircraft,” Journalof aircraft, Vol. 47, 2010, pp. 956 – 970.

[4] van Dommelen, J. L., “Design of a Forward-Swept Blended Wing Body Aircraft,” Tech.rep., Delft University of Technology, Oct. 2011.

[5] Roskam, J., Airplane design, Part I: Prelim-inary sizing of airplanes, The University ofKansas, 1990.

[6] Roskam, J., Airplane Design Part I: Prelimi-nary Sizing of Aircraft, DARcorp, Lawrence,KS, 2005.

[7] Morris, A. J., “MOB A European DistributedMulti-Disciplinary Design and OptimisationProject,” Tech. Rep. AIAA 2002-5444, Cran-field University, Sep. 2002.

[8] Raymer, D., Aircraft design: A conceptual ap-proach, American Institute of Aeronautics andAstronautics, Inc, 2006.

[9] Roskam, J., Airplane design, Part VI: Pre-liminary calculation of aerodynamic, thustand power characteristics, The University ofKansas, 1990.

[10] Kroo, I., Aircraft design, Synthesis and Analy-sis, Desktop Aeronautics, Inc, Jan 2001.

[11] Torenbeek, E., “Development and Applic-cation of a Comprehensive, Design-sensitiveWeight Prediction Method for Wing Structuresof Transport Category Aircraft,” Tech. rep.,Delft University of Technology, Department ofAerospace Engineering, 1992.

[12] F. J. J. M. M. Geuskens, O. K. Bergsma, S. K.and Beukers, A., “Analysis of ConformablePressure Vessels: Introducing the Multibubble,”Tech. Rep. AIAA 54561-869, Delft Universityof Technology, Aug. 2011.

[13] Obert, E., Aerodynamic design of transport air-craft lecture notes, Delft University of Technol-ogy, 2007.

[14] Rocca, G. L., Langen, T., and Brouwers, Y.,“The Design and Engineering Engine: Towardsa Modular System for Collaborative AircraftDesign (accepted),” ICAS 2012, Brisbane, Aus-tralia, September 2012.

Nomenclature

b Wing span, mc Chord, mCnβ Weathercock stability, 1/radClβ Effective dihedral, 1/radClmax Maximum lift coefficient, -Mcruise Cruise Mach number, -l Aircraft overall length, mL/D Lift-to-drag ratio, -s Runway length, mScabin Cabin floor area, m2

Vcargo Cargo volume, m3

Vrot Take-off rotation speed, m/sVstall Stall speed, m/sWf Fuel weight, kgWOE Operational empty weight, kgWpl Payload weight, kgWTO Take-off weight, kgx Aircraft x-position, mδ Control surface deflection, degε Wing twist, degγ Climb gradient, %Γ Wing dihedral, degΛ Wing sweep, deg

12

A Conceptual Design and Analysis Tool for Blended Wing Body Aircraft

Subscripts

0 zero altitudeclean clean configurationland landingTO take-off

Abbreviations

BWB Blended Wing BodyCS Certification SpecificationDD drag divergenceFAR Federal Aviation Regulationsle leading edgeMAC Mean aerodynamic chordmlg main landing gearnlg nose landing gearOEI one engine inoperativeSM static marginTAW Tube and Wing Aircraftte trailing edgeTO take-off

The authors confirm that they, and/or their com-pany or organization, hold copyright on all of the orig-inal material included in this paper. The authors alsoconfirm that they have obtained permission, from thecopyright holder of any third party material includedin this paper, to publish it as part of their paper. Theauthors confirm that they give permission, or have ob-tained permission from the copyright holder of thispaper, for the publication and distribution of this paperas part of the ICAS2012 proceedings or as individualoff-prints from the proceedings.

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