design and test of a uav blended wing body ... - icas

DESIGN AND TEST OF A UAV BLENDED WING BODYCONFIGURATION

Kai Lehmkuehler∗ , KC Wong∗ and Dries Verstraete∗∗School of Aerospace, Mechanical and Mechatronic Engineering, The University of Sydney,

[email protected]

Keywords: Blended Wing Body, UAV, Wind tunnel, Stability, Propulsion Effects

Abstract

This paper presents the design and test of a UAVblended wing body configuration. It is the re-sult of a global student design project betweenthe University of Sydney, the University of Col-orado and the University of Stuttgart. Firstly,the design methodology and constraints are in-troduced. Then the wind tunnel test setup is de-scribed and the data presented. Finally, an engi-neering method to predict the propulsion effectson this unusual airframe is described.

Comparison between the wind tunnel dataand panel code predictions shows good agree-ment, also reinforced by successful flight tests.

Propulsion effects on a low stability airframecan be serious and need to be considered duringthe design to avoid possible instabilities. Themethod presented allows for a quick estimatewithout tedious computations or tests.

1 Introduction

Blended wing body (BWB) configurations haveattracted considerable interest lately as an alter-native, more efficient platform for transport air-craft [1][2]. In conventional tube-wing config-urations the fuselage typically contributes onlyminor amounts of lift while adding considerableskin friction drag as well as a disruption of thelift distribution across the wing. BWBs locatetheir payload volume inside the wing such thatthe entire external surface contributes to the gen-

eration of lift and the aircraft can be shaped foran optimal lift distribution. Typically they alsoare tailless flying wings which potentially furtherreduces the drag.

The use of BWBs for full scale transport air-craft is still some time away due to several prob-lems; mainly the compliance with current reg-ulations is difficult if not impossible (passengerevacuation, aircraft control without artificial sta-bility and so on). Another application for a BWBmight be a smaller, unmanned platform, wheremost of these constraints do not apply but highefficiency and large internal volume are required.Therefore the aim of this project is to design andtest a small scale BWB and to determine if theplatform is a viable configuration as a UAV.

The aircraft discussed in this paper is the firstiteration of the project. It was designed for theinternational Hyperion project, which was a co-operation of student teams from Sydney, Aus-tralia, Stuttgart, Germany and Colorado, USA.The Sydney team was tasked with the airframeconceptual design and the wind tunnel testing.The plane was built in Germany and the US,where it was flown successfully in April 2011.

This paper will outline the design and testingconducted before the first flight. Future publica-tions will cover the flight testing and the continu-ing development of the platform.

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KAI LEHMKUEHLER, KC WONG AND DRIES VERSTRAETE

2 Design Methodology

There was a constraint of 9 calender months totake the project from requirement definition tofirst flight. Thus time was limited for the concep-tual design phase. Subsequently, the focus wasput on stability and control of the aircraft, withan optimisation for minimum drag to follow in alater stage.

A tailless aircraft has some unique require-ments for stable flight especially in the pitch axis.The moment arm between the CG and the eleva-tor is short and the design is a compromise be-tween the static margin and the elevator effective-ness. The two extremes are:

1. A forward neutral point (NP) gives goodelevator effectiveness but it will be difficultto place the CG sufficiently forward of theNP to ensure pitch stability.

2. An aft NP allows for a more practical CGplacement but the elevator moment arm be-comes shorter which makes trim difficultwith sensible elevator sizes. In this ar-rangement it is also difficult to achieve ap-propriate pitch damping.

The preliminary design was done in AVL [3],which allows for quick evaluations of differentairframe shapes with immediate output of all sta-bility derivatives. The design was later investi-gated with PanAir [4], a fully 3-d panel code, andthe results compared to AVL and the wind tunneldata.

The basic parameters for the aircraft were:16kg MTOW, 3m span and 15m/s stall speed.Together with the general requirements of a fly-ing wing mentioned above, this placed some verydifficult constraints on the design, especially forthe test flights in Colorado which is at an alti-tude of 1600m above sea level. A BWB cannothave flaps due to the additional pitching moment.Hence the clean airframe had to produce a maxi-mum lift coefficient of at least 0.8, which is morethan has been reported by other designs of thistype [5]. The constraint on the wing span limitedthe aspect ratio and therefore the maximum wingarea possible.

Fig. 1 PanAir model with pressure distributionand surface flow directions

All BWBs currently under investigation in lit-erature have highly swept wings for their tran-sonic flight regime. As this aircraft is flying sub-sonic at all times the wing sweep was limited to15 degrees (for roll stability) to avoid tip stall ten-dencies associated with higher sweep at high liftcoefficients.

As for any flying wing, the aerofoils used re-quire a very low moment coefficient, which typi-cally leads to reflexed sections. For a small UAVthe low flight speed and wing chord also requiresgood performance at low Reynolds numbers. Theaerofoils for the plane were selected based onthe excellent website for model aircraft builders[6], which provides comprehensive informationon low Reynolds number, low pitching momentaerofoils. The chosen section was the S5010 [7]aerofoil, which is 10% thick. For the body sec-tion, the aerofoil was modified in X-Foil [8] intoa S5016 with 16% thickness to provide the re-quired volume.

The design process for the shape of the air-craft was basically a manual exploration of thedesign space with several configurations andtrends examined by hand until an arrangementwas found to fit the requirements. A formal op-timisation process was not possible due to thetime constraints. Further work since then has de-termined that within the requirements and con-straints there is no significant improvement pos-sible unless the wingspan constraint was relaxed.

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Blended Wing Body UAV

The final airframe shown in Figures 1 and12 features a tractor propeller in the nose, whichwas chosen over an initial pusher concept mainlybecause of clearance issues during rotation andweight and balance advantages. The propeller isalso more efficient when not running in the flowfield of the main body. Other configuration deci-sions include:

• A single large elevator 20% of the mainbody chord to ensure enough control powerfor all flight phases;

• Twin fins on each side of the elevator thesize of which was determined from the sug-gested values for the weathercock stabilityCnβ in [9];

• Upwards canted swept back wing tips foradditional dihedral;

• Tricycle landing gear; and

• The final wing area being 1.53 sqm with abody length of 1.25m.

3 Wind Tunnel Set-up

The department’s 7x5 feet low speed wind tun-nel was the ideal venue for testing this airframe,as it could accommodate a half-scale model at therequired speed range to match the Reynolds num-bers between the test and the flight airframe.

To perform the tests for this aircraft, the fullflight envelope in angle of attack and sideslip wasrequired. The existing balance could move onlyin angle of attack with the model suspended fromthe roof. It was thus required to design a new bal-ance using the existing turntable mechanism. Thedesign chosen was a single sting balance with aninternal loadcell as shown in Figure 2. The modelis driven by an actuator in angle of attack on topof the sting and the entire mechanism rotates insideslip on the turntable. The loadcell process-ing and the motion control is handled in a newlydeveloped graphical MatLab application, whichallows control of all features from a single GUI.An image of the setup is shown in Figure 3.

Fig. 2 CAD model of the new 6 axis wind tunnelbalance

The model was constructed using the recentin-house completed CNC hot wire cutter, whichwas built by the UAV design group. The machineallows to cut linearly tapered foam blocks withgood precision but it cannot generate the curva-tures required for the blended shape. These werefinished by hand after the assembly of the blocksinto the complete airframe. This method of con-struction allows for a reasonably quick modelconstruction on a limited budget but the final re-sult has inevitably some tolerances due to thesmall scale and manual re-work. This was toshow up later in the wind tunnel results but wereconsidered acceptable for this first demonstratorairframe. A second limitation for the testing isthe lack of stiffness of the available load cell inthe roll axis. Hence the model vibrates slightly inroll, especially near stall, where the vortex shed-ding from the wing tips becomes more severe. Asexpected, the roll moment data shows more errorthan the other axes but the data was still usable asdiscussed below.

The model features an interchangeable nosesection for powered testing and removable wingtips to evaluate different designs.

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Fig. 3 Wind tunnel installation with fairings removed

4 Test Results for the Unpowered Airframe

This section presents the test results for the half-scale airframe model using the new balance. Thedata represents the current status of maturing ex-perimental setup. Thus some minor errors areexpected to be in the data. The data has beencorrected for the changing gravity vector duringangle of attack changes but no corrections havebeen applied for wind tunnel effects like block-age. Those are still under investigation.

All data is reported at 20m/s tunnel speed,which represents a Reynolds number of 450,000based on the mean aerodynamic chord. Whereapplicable, changes due to different airspeeds arediscussed. On all plots blue dots represents theexperimental data points, solid red the appropri-ate fit and dashed black the data from PanAir. Asa comparison, data from AVL has been included,too to judge the performance of this code againstPanAir for an unusual airframe concept.

4.1 Lift

The lift curve shown in Figure 4 matches the pre-dictions closely. The error is less than 2% at20m/s. At lower speeds the difference reduces(no error at 13m/s) and it increases with airspeed.The aircraft has a relatively low aspect ratio of

5.9, so the slope of the lift curve is low as ex-pected.

CLα experiment 4.15 / radCLα PanAir 4.23 / radCLα AVL 4.016 / radCL,max experiment 0.946

−4 −2 0 2 4 6 8 10 12 14 16−0.4

−0.2

0

0.2

0.4

0.6

0.8

1

1.2

α [deg]

CL

Fig. 4 Lift coefficient at 20 m/s

The maximum lift coefficient increases withairspeed to 0.96 from 0.93 at 13m/s. This iscaused by the increase in Reynolds number overthe wings.

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The stall pattern is well behaved with flowseparation starting at the inboard wing joint aspredicted. The wing tips with the sharp break inthe leading edge stall at about the same time butthe flow separation does not spread inboard. Theshape of the wing tips is currently under revisionfor that reason.

4.2 Drag

The drag polar in Figure 5 shows a small shiftbetween the experiment and the predictions. Thewind tunnel data has the minimum drag at CL ≈0.1 compared to CL≈ 0.05 from PanAir. This canbe caused by small differences in the geometry ofthe model to the PanAir geometry.

The span efficiency is generally lower thanthe inviscid predictions up until 30m/s, where itthen matches. This is caused by the low Reynoldsnumbers, which are not taken into account by thepanel methods. The AVL prediction for the spanefficiency is very optimistic, with the differencecaused mainly by small differences compared toPanAir around the zero-lift drag. This leads tolarge changes in the quadratic fit used to find e.

Span efficiency e experiment 0.72Span efficiency e PanAir 0.77Span efficiency e AVL 0.98CD,0 experiment 0.015

As mentioned before, this aircraft was mainlydesigned for stability and control properties withonly limited effort on optimising the flight per-formance. The high induced drag has since thenbeen investigated and identified to be caused bythe sudden, large change in chord in the body-wing transition and the low aspect ratio. Thesefindings will be used for the next iteration of thevehicle design.

A model aeroplane style fixed landing gearhas also been tested in the wind tunnel. Onthis scale a landing gear is usually large com-pared to the rest of the airframe dimensions todeal with rough landing strips used to fly theseplanes. Here the landing gear adds about 30% tothe zero lift drag. Clearly, this eliminates any per-

−0.4 −0.2 0 0.2 0.4 0.6 0.8 10

0.05

0.1

0.15

0.2

0.25

CL

CD

Fig. 5 Drag polar at 20 m/s

−4 −2 0 2 4 6 8 10 12 14 16−15

−10

−5

0

5

10

15

20

α [deg]

L/D

(in

cl. C

D0)

Fig. 6 L/D at 20 m/s

formance gain of the blended airframe and shouldbe avoided.

The combination of low aspect ratio and highinduced drag results in a maximum L/D of 19.3as shown in Figure 6. While this should improvein future versions of the airframe, it is still a re-spectable value for a low aspect ratio aeroplaneat low Reynolds numbers.

4.3 Pitching Moment

There is a 8% error in the pitching moment slopeas shown in Figure 7. This is mainly caused bythe limited accuracy of the model (especially thereflexed aerofoils on the wings) and also someuncertainty in the location of the mounting point

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for the loadcell. A change in CG position of only3mm eliminates the error completely.

The CG range of the aircraft is very small,with only 30mm between instability and the lim-its of the elevator effectiveness. The flight CGhas been placed at x = 0.6m based on pilot com-ments.

The landing gear tests revealed only negligi-ble increases in nose down pitching moment dueto the additional drag of the gear.

Cmα experiment -0.57 / radCmα PanAir -0.62 / radCmα AVL -0.564 / radNP experiment 0.646 m

−4 −2 0 2 4 6 8 10 12 14 16−0.2

−0.15

−0.1

−0.05

0

0.05

α [deg]

Cm

Fig. 7 Pitching moment at 20m/s with Xre f = 0.55m

4.4 Side force

The side force derivative shows good agreementbetween predictions and the experiment.

Cyβ experiment -0.1863 / radCyβ PanAir -0.1835 / radCyβ AVL -0.187 / rad

4.5 Rolling moment

The experimental data for the rolling moment inFigure 8 shows more fluctuations than the otheraxes. This is due to the vibrations of the model

about the roll axis on the loadcell as discussedbefore.

The value of the derivative matches the AVLprediction quite well but there is a large error tothe PanAir computation. The reason for this errorhas not yet been determined.

Clβ experiment -0.059 / radClβ PanAir -0.0258 / radClβ AVL -0.059 / rad

−4 −2 0 2 4 6 8 10−10

−8

−6

−4

−2

0

2x 10

−3

β [deg]

Cl

Clb

= −0.0590

Clb

(PanAir)= −0.0258

Fig. 8 Rolling moment at 20m/s

The airframe has no dihedral for ease of con-struction and was designed to obtain its roll sta-bility from the limited wing sweep and the body-wing flow field. The upward canted wing tipsprovide further stability but this is not really re-quired, based on the first flight tests.

4.6 Yawing moment

The yawing moment shown in Figure 9 is slightlystiffer than predicted but the derivative is still be-low the recommended value of Cnβ = 0.05 [9].This can be easily fixed by slightly larger fins butagain has been found to be not necessary duringthe initial flight tests.

Cnβ experiment 0.0365 / radCnβ PanAir 0.033 / radCnβ AVL 0.03445 / rad

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−4 −2 0 2 4 6 8 10−3

−2

−1

0

1

2

3

4

5

6

7x 10

−3

β [deg]

Cn

Cn

b

= 0.0365

Cn

b

(PanAir)= 0.0330

Fig. 9 Yawing moment at 20m/s

4.7 Pilot feedback

The airframe has been flown by several pilots,full (Figure 10) and half-scale versions in theU.S.A. and a half-scale version in Australia. Pi-lots’ feedback on the handling qualities reportedgood characteristics in flight and some difficultieson take-off and landing. In flight, the aircraft wasreported quite stable in all three axes with someminor adverse yaw. It was well controllable inall manoeuvres tested with well damped modesof motion. The amount of elevator to trim is rela-tively large but this is expected for a flying wingconfiguration.

On take-off, a reduced longitudinal stabilitywas observed due to propulsion effects as dis-cussed below. This was manageable in the U.S.A.as the flights there were done on a sealed run-way and therefore very limited disturbances dur-ing the critical speed range. In Australia, there isno convenient access to such a runway, so grassfields were used. Here the natural unevenness ofthe ground has caused more severe issues duringthe take-off run with several premature lift-offsbelow rotation speed.

On landing, the reduced pitch damping makesit difficult to flare due to phugoid oscillations.This issue can be improved with better pilot train-ing but it is intended that subsequent designs willimprove on these undesired handling qualities.

Fig. 10 First flight of the full scale HyperionBWB in Boulder, Colorado

5 Propulsion Effects

Any airframe shows some effect of its propul-sion system on its stability characteristics [10].For example, on a conventional, propeller-drivengeneral aviation aircraft, the swirl of the propwash impacting the vertical stabilizer causes ayawing moment. Likewise, on a jet with under-wing engines there is a thrust dependent pitch upeffect.

This BWB configuration with a tractor pro-peller in the nose shows an additional behaviour.The propeller is effectively blowing the inboardsections of the body behind it, causing a locallydifferent dynamic pressure over these aerofoils.This thrust and airspeed dependent effect changesthe moment balance between the the body andthe outer wings. As the body creates more liftin the slipstream compared to the non-poweredairframe, its more forward neutral point becomesmore dominant and the overall NP moves for-ward. With the limited CG range of the config-uration, this may have a severe effect, as the NPcould potentially move in front of the CG duringlow airspeed, full power flight conditions, caus-ing pitch instability.

With the limited time frame of the projectspent predominantly with the basic design and

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testing, this effect was not investigated untilshortly before the first flight. After it becameapparent that this might have catastrophic effectsduring the take-off run, a quick method to esti-mate the magnitude of the NP movement was re-quired. Further wind tunnel work was impossibleat that stage and there is no quick way of simu-lating these non-uniform inflows in the analysistool used for this project. Thus a method was de-vised using the PanAir results, which is presentedbelow.

PanAir reports all forces and moments for theentire airframe and also for each parallel, span-wise panel strip (Figure 12) from the leadingedge to the trailing edge. Instead of calculatingthe NP from the full aircraft data with

X̄NP = X̄CG−Cmα

CLα

(1)

it is possible to calculate the NP of every stripand find the overall NP by a weighted average ofall strips. Then, using the velocity profile behinda propeller from a blade element code, each stripcan be adjusted to the local dynamic pressure tosimulate the effect of the prop-wash. The steps indetail:

1. Determine the slipstream velocity distribu-tion for the propeller with a blade elementcode. The propeller for the aircraft had 4-blades and 20 inch diameter. As the aero-foils on the model aircraft propellers areusually not known, a simple Clark Y sec-tion was run in X-foil at the 75% stationat full power. Here the propeller speedwas 8000 RPM, giving a Reynolds num-ber of 230000 at Mach 0.5. The lift curveand drag polar at that flow condition weredetermined and used in the blade elementcode. Tip losses were added by imposingzero velocity at the propeller tips. The pro-peller had 10 inch geometric twist alongthe blade.

2. Using this data shown in Figure 11, the lo-cal dynamic pressure seen by every stripcan be interpolated.

0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 120

25

30

35

Prop Station

Slip

stre

am V

el [

m/s

]

Fig. 11 Slipstream velocities for full power atV∞ = 20m/s

3. Dimensionalise the lift and pitching mo-ment of every strip using the local dynamicpressure to obtain the actual forces and mo-ments acting on each strip.

4. Determine the NP of every strip using eq.1. This requires a PanAir solution at twoangles of attack to determine the lift- andmoment curve slopes.

5. Calculate the NP of the aircraft by apply-ing a weighted average for each strip basedon the amount of lift produced compared tothe total lift.

6. Compare the result to the NP of the unpow-ered aircraft for several airspeeds.

After adjusting the NP of the PanAir resultsto match the wind tunnel results by a small ref-erence point change as mentioned above, a testcase with zero slipstream yields Xnp = 0.645mvs. Xnp = 0.644m for the full aircraft calculation.This small difference is probably caused by in-ternal round off errors as the coefficients are onlyavailable to 5 digits in the PanAir solution file.A further source of error might be the weightedaveraging based on the lift produced per strip.However, the difference is small for practical use,so no further improvements were attempted.

Crucial for the method to predict useful re-sults is the number of panel strips affected by the

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Fig. 12 Propeller slipstream distribution (As-sumption blue (top), Experiment green (bottom)

slipstream. Several graphs of propeller body in-terference are given in [11]. There it can be seenthat the typical contraction of the slipstream be-hind the propeller disc does not happen due tothe presence of the body blockage. Instead, theslipstream follows the contour of the body. Toestimate the affected area of the BWB this datawas combined with some flow visualisation doneon the un-powered model in the wind tunnel be-fore. It was assumed that the slipstream followsthe contour of the body up to the impact with thewing as shown in the top half of Figure 12 in blue.Over the wing it would then align with the freestream. This distribution results in 12 strips, or1.9 times the prop diameter, being affected by theslipstream and hence the blade element resultswere interpolated over these 12 strips as shownin Figure 12.

5 10 15 20 25 30 35

0.58

0.59

0.6

0.61

0.62

0.63

0.64

0.65

0.66

Speed [m/s]

Xnp

[m]

Experiment half scale 450WExperiment windmillingPanAir prediction 10 stripsPanAir prediction 12 strips

Fig. 13 NP changes due to propulsion effects (In-dicated are the stall speed and flight CG)

The result of these computations show a NPshift forward during slow, high power conditionsfrom Xnp = 0.645m to Xnp = 0.61m at stall speed.That means that the static margin just before ro-tation is reduced to 1.4%, down from 6.4% atcruise. This is clearly a severe effect and wasobserved during the take off runs preceding firstflight with the aircraft pitching up uncontrollablyduring acceleration. Subsequently, based on thisanalysis, the CG was placed further forward andthe flight was a success.

6 Test Results for the Powered Airframe

To verify the results, powered tests were per-formed using the half- scale wind tunnel model.It was powered by a 550W motor driving a10 inch propeller with two blades. To achievereasonably similar conditions, the test airspeedswere scaled by l0.5 and the power was scaledas l3.5, from 5000W on the full scale aircraft to450W. The data is plotted in Figure 13, scaledto the experiment speeds. To start with, it wasinvestigated if the unpowered, windmilling pro-peller had any influence on the NP (green). It canbe seen that at low speed the effect is negligible,but at higher airspeeds there is a small shift for-ward.

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The results of the powered experiment areplotted in black. They match the computations(cyan) well around the stall speed (14% differ-ence in NP shift) but diverge at high speeds. Herethe real propeller is still generating thrust whilethe blade element code shows windmilling from18m/s. This is caused by inaccuracies in thepropulsion model, which amongst others ignoresthe effects of the body blockage and changesto the flow direction through the propeller disc.Improvements on the modelling of these effectsshould increase the accuracy of the simulation,but this was impractical for the simplistic ap-proach taken.

The assumed flow field over the aircraft wasinvestigated on the model with tufts and a raketracing the propeller tip vortex. The result isshown in Figure 14 with the blue string and hasbeen indicated on Figure 12 in the lower half.The stream tube right behind the prop contractsthe usual way contrary to the assumption. Thehighly tapered body does not produce enoughblockage to match the results from [11], whichused a blunter body. Further down the the flowaligns with the body contour as expected and thenit curves in over the wing leading edge. From10% chord the slipstream is aligned with the freestream to the trailing edge. The measured dis-tance from the centreline is 190mm, which trans-lates into 10 strips or 1.5 times the propeller di-ameter on the PanAir model in Figure 12. There-fore, the affected portion of the airframe wasslightly overestimated. A calculation with 10strips is shown in red in Figure 13. The NP shiftbecomes more severe at low speeds as the slip-stream is concentrated over the inboard parts ofthe BWB. This means that the propulsion modelproduces too much thrust at these low speeds,which will be caused by the simplifications men-tioned above. Again, a better model should im-prove the accuracy.

The presented method shows a simplified, butfast way of determining the propulsion effect onthis BWB airframe. It was very helpful under-standing the effect and it is conservative as itover-predicts the effect by a reasonable amount.The results of the method were crucial for the

Fig. 14 Prop wash boundary as determined in thewind tunnel for a 10 inch propeller

success of the first flight and will be tested onfuture versions of the aircraft with different bodyshapes to explore its versatility.

7 Conclusion and Future Work

This paper presented a design and testing ofa blended wing body UAV airframe. The de-sign methodology using fast panel methods hasbeen proven viable for an unusual configuration.The wind tunnel tests matched the predicted datawell and the flight testing revealed good handlingqualities in flight. Some problems during takeoff and landing due to the limited aircraft stabil-ity and the presence of propulsion effects on thelongitudinal stability remain. The method usedto obtain an engineering estimate of these effectshas been proven usable.

Future development of the airframe will fo-cus on improving the induced drag properties andthe remaining stability issues. A full flight instru-mentation system is currently being developedto perform flight testing for parameter estimationpurposes.

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References

[1] Hileman, J. I., Z. S. Spakovszky, et al. (2007)."Airframe Design for Silent Aircraft." 45thAIAA Aerospace Sciences Meeting and Exhibit8 - 11 January 2007, Reno, Nevada.

[2] Liebeck, R. H. (2004). "Design of the BlendedWing Body Subsonic Transport." Journal ofAircraft 41(1).

[3] Drela M. (2011). AVL. Retrieved 15.3., 2012,from web.mit.edu/drela/Public/web/avl/

[4] Carmichael M. (2010). "Panair" Retrieved 15.3.,2012, from www.pdas.com

[5] Scholz, D. (2007). "A Student Project of aBlended Wing Body Aircraft: From Concep-tual Design to Flight Testing." Presentation forEWADE 2007.

[6] Siegmann, H. (2008). "Aerodesign (inGerman)." Retrieved 15.3., 2012, fromwww.aerodesign.de.

[7] University of Illinois. "UIUC Airfoil Co-ordinates Database." Retrieved 15.3., 2012,from www.ae.illinois.edu/m-selig/ads/coord-database.html.

[8] Drela M. (2011). X-foil. Retrieved 15.3., 2012,from web.mit.edu/drela/Public/web/xfoil/

[9] Raymer D. (2005). Aircraft Design - A concep-tual approach, AIAA

[10] Nelson, R. C. (1998). Flight Stability and Auto-matic Control, McGraw Hill.

[11] C.N.H. Lock, M. A. (1929). "Airscrew BodyInterference: An Examination of a Method ofCalculating the Mutual Effect of Airscrew andBody by the Strip Theory." Aircraft Engineer-ing and Aerospace Technology, Vol. 1 Iss: 6,pp.207 - 209.

7.1 Copyright Statement

The authors confirm that they, and/or their company ororganization, hold copyright on all of the original ma-terial included in this paper. The authors also confirmthat they have obtained permission, from the copy-right holder of any third party material included in thispaper, to publish it as part of their paper. The authorsconfirm that they give permission, or have obtainedpermission from the copyright holder of this paper, forthe publication and distribution of this paper as part of

the ICAS2012 proceedings or as individual off-printsfrom the proceedings.

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