airfoil selection of mav (miniature air vehicle)

7/18/2019 Airfoil Selection of MAV (Miniature Air Vehicle)

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ISSN : 2319 – 3182, Volume-2, Issue-4, 2013

38

Airfoil Selection of MAV (Miniature Air Vehicle)

for Low Reynolds Number

Mayur S. Marathe & S. N. Bansode

K.J Somaiya College of Engineering, Mumbai University

E-mail : [email protected] [email protected]

Abstract – This paper discusses issues and practical

requirements of Airfoil for MAV. Here considering the

MAV which travel with the speed range between 9-20 m/s.The Airfoil which is been selected on various criteria, i.e. -

stable flight, cover maximum distance with minimum

force. So here the NACA 2204 is been selected for MAV.

The Fluent analysis is done on the airfoil for lift to drag

ratio. These MAV are having some purpose i.e.:- they can

be use as a spy in enemy area, inspection of hazardous

area, where human resource can’t reach. Aerodynamic

performance and stability should be considered in the

context of the airfoil structural integrity. Particular

attention should be paid to the unsteady nature of the flow.

Keywords – MAV, NACA 2204, Li ft to Drag ratio, Fluent

analysis.

I. INTRODUCTION

Airfoil for MAV is very important, as it has to

travel comparatively more distance & stable flight for

the given forces. Here NACA 2204 is been selected as

it’s a low camber airfoil. The lower surface of airfoil is

somewhat flat, so it doesn’t allow the air flow away

from it when it glides down to the surface. The air that

hits to the lower surface of an airfoil try to push or lift

the MAV, as the MAV is coming down the forces the

resultant force will be in downward direction, so it willglide down to surface comparatively at slow speed,

which will cause minimum damage. At the same time

the airfoil also has to be good lifting coefficient i.e. Liftto drag ratio has to be high. As the MAV has to attain

the height within in short range of distance, the stalling

angle of airfoil also has to be high.

Fig. 1: Airfoil with angle of attack

II. DEVELOPMENT CONSIDERATIONS

Several areas need to be carefully considered for the

selection of a practical airfoil, including aerodynamics.These will be covered in turn, following a discussion of

the benefits of airfoil. Consider the simple wing

geometry as shown in figure 1 [1]

This geometry will be used throughout the

remainder of this paper as a baseline. Here the airfoil is

at α angle of attack with relative wind (V∞).

III. EFFICIENCY

For a selected airfoil, we are principally interested

in maximizing lift L and minimizing the drag D, or

alternatively, maximizing the lift-to-drag ratio, L_D

(also written as the ratio of lift coefficient (C_l) to dragcoefficient (C_d) or C_l /C_d, defined below). It is also

necessary to look out on the overall efficiency of a wing.

This ratio depends on wing geometry & air flow

condition. These flow conditions are expressed as

dimensionless parameters such as the Reynolds numberRe and Mach number M. A selected airfoil profile will

have vastly different lift and drag characteristics over

the possible ranges of Re and M for a profile selected.

Thus, airfoils are typically designed for a narrow range

of flow conditions for optimum performance.

Alternatively, one could design an airfoil that willoperate over a wide range of air flow conditions.

Lift capability & drag capability of an airfoil is

depended on Lift & Drag coefficient, which is given as

follows [1]

L=0.5*ρ*V^2*S*C_l

D=0.5*ρ*V^2*S*C_d

Furthermore, the total drag is further subdivide into

number of drag, such as, form, pressure, skin friction,

parasitic, induced & wave drag. The induced drag can

be estimated in terms of wing geometry by



International Journal on Theoretical and Applied Research in Mechanical Engineering (IJTARME)


39

D_i= (C_l^2)/πAR

In steady & level flight the Drag force has to be

balance by thrust & weight must be balance by Liftforce.

Fig 2 : Level flight condition

T=D

L=WAbove quation conclude that as we increase the lift

i.e.:- C_l, the weight carrying capacity increase. For

increasing the lift to drag ratio, drag has to reduce, as thedrag decrease, the thrust require to lift the flight is

comparatively less, due to which the flight efficiency

increases & hence the performance also increase. Hence,

it can be say that with increase the lift to drag ratio, the

efficiency of the flight increase

L/D=C_l/C_d

Lift & drag also help to reduce the gliding angle,

lesser the gliding angle more will be the distance travel

in horizontal direction & vice-versa. If a glider is in asteady (constant velocity and no acceleration) descent, it

loses altitude as it travels. The glider's flight path is a

simple straight line, shown as the inclined red line in the

figure. The flight path intersects the ground at an angle

called the glide angle [2]. If the flown distance (d) is

known to us and the altitude change h is also known to

us, the glide angle can be easily calculated using

trigonometry.

Fig 3: Gilding angle

From above fig 3 it can be easily conclude that lesser the

gliding angle more will be the distance travel.

IV. AERODYNAMICS

Below fig will compare the symmetric & non-

symmetric (cambered) airfoil. The air is flown over theairfoil at different angle of attack [4]. The symmetric

airfoil stall at greater angle of attack as compare tocambered airfoil. Even at zero angle of attack, cambered

airfoil will generate some positive lift coefficient, but in

symmetric it produces zero lift coefficients.

Fig 4: Comparison between cambered & symmetric

airfoil

From above fig 4 it can be say that cambered airfoil

has a lower stall angle than the symmetric one. Thus

from here it can be conclude that cambered airfoil isused.

As the airfoil is selected, the gliding of the flight

should be good. So airfoil has to be selected in such away that has less cambered as compared to other

airfoils, i.e.: the lower surface of an airfoil has to benearly equal to flat. Reason for such airfoil is that the

flat surface doesn’t allow the air to flow away from the

edges of an airfoil, due to which italways try to lift the

flight against the gravity, which indirectly increase lift

to drag ratio & hence the efficiency. Form above reason

NACA-2204 is chosen, which satisfy requirements.

Fig 5: NACA-2204

V. CFD ANALYSIS OF AIRFOIL

Here NACA-2204 airfoil is selected, so here the

CFD analysis is done on the airfoil so that the lift to

drag ratio is determine at various angle of attack. CFD

analysis is done in ANSYS. For analysis, the CAD

model is prepared in CATIA or AUTOCAD or anyother CAD software. These CAD model is import in





40

meshing software, here the meshing is done in ICEM.

Below fig6 give the view of mesh airfoil.

Fig 6 :Mesh of an airfoil

Accuracy of the result depends on the quality ofmesh that has made; high quality of mesh will give good

results. This mesh airfoil is imported in ANSYS fluent

for fluid analysis. Here the operating condition will be

as per the requirements, i.e.:- the velocity should be 9 -

20 m/s, angle of attack should vary from -5° to 5°. Herethe in-viscid flow is considered, as air is for MAV in-

compressible flow, it is travelling at lower altitude

VI. FLUENT ASSUMPTIONS

Fluent calculation is done on the 2D airfoil with in-

viscid, in-compressible flow & with no shock waves. It

solves the conservation of energy, momentum & massacross the grid of an airfoil with Mach number ranging

from 0.026 – 0. 054. The grid used for this analysis is

having coarse mesh around the inlet, outlet & boundary

& the grid is dense around the airfoil, as we can see in

above fig 6.

VII.REALIZABLE K- Ε SOLUTION PROCEDURE

The Realizable k- ε solution procedure is an

essential part of the present design/analysis method. The

term ``realizable'' means that the model satisfies certain

mathematical constraints on the normal stresses,

consistent with the physics of turbulent flows. Tounderstand this, consider combining the Boussinesq

relationship and the eddy viscosity definition to obtain

the following expression for the normal Reynolds stress

in an incompressible strained mean flow

The weakness of the standard - model or other

traditional - models lies with the modeled equation for

the dissipation rate (ε). The well-known round-jet

anomaly (named based on the finding that the spreading

rate in planar jets is predicted reasonably well, but

prediction of the spreading rate for axis-symmetric jets

is unexpectedly poor) is considered to be mainly due to

the modeled dissipation equation

VIII. BOUNDARY CONDITIONS

Inlet: - Velocity inlet boundary conditions are used to

define the flow velocity, along with all relevant scalar

properties of the flow, at flow inlets. The total (or

stagnation) properties of the flow are not fixed, so theywill rise to whatever value is necessary to provide the

prescribed velocity distribution.

This boundary condition is intended for

incompressible flows, and its use in compressible flows

will lead to a nonphysical result because it allows

stagnation conditions to float to any level. You should

also be careful not to place a velocity inlet too close to asolid obstruction, since this could cause the inflow

stagnation properties to become highly non-uniform.

Outlet: - Pressure outlet boundary conditions require the

specification of a static (gauge) pressure at the outlet

boundary. The value of static pressure specified is usedonly while the flow is subsonic. Should the flow become

locally supersonic, the specified pressure is no longer

used; pressure will be extrapolated from the flow in the

interior. All other flow quantities are extrapolated from

the interior.

A set of ``backflow'' conditions is also specified to

be used if the flow reverses direction at the pressure

outlet boundary during the solution process.

Convergence difficulties will be minimized if you

specify realistic values for the backflow quantities

Wall: - Wall boundary conditions are used to bound

fluid and solid regions. In viscous flows, the no-slip

boundary condition is enforced at walls by default, but

you can specify a tangential velocity component in

terms of the translational or rotational motion of the wall

boundary, or model a ``slip'' wall by specifying shear.

IX. RESULTS

The results presented here are aimed to select the

best airfoil for MAV. The polar graphs were calculated

by specifying sequence of angle of attack in increment

of 1 degree. Since a good initial guess was available for

each point from the previous angle of attack, realizable

required at least 10000 iteration to coverage

solution.

Below graph 1 will show the Lift versus angle ofattack for FLUENT. The angle of attack is varied from -

5° to 5° with 1° of interval.





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Graph 1 : C_l vs α

Table 1: Values of C_l varies with α

Below graph 2 will show the Drag versus angle of

attack for FLUENT. The angle of attack is varied from -

5° to 5° with 1° of interval.

Graph 2 : C_d vs α graph

Table 2: Values of C_d varies with α

As in this paper the main concentration is the lift to

drag ratio, so below graph 3 will give the clear view that

how does C_l/C_d varies with respect to angle of attack.From below graph it can be conclude that NACA-2204

is having maximum C_l/C_d at 3° angle of attack.

Graph 3: C_l /C_d vs α graph

Table 3: Values of C_ l/C_d varies with α

Pressure distribution of airfoil is also factor for

determination of lift. In this paper NACA- 2204 is

having very less cambered due to which it has somewhat

low surface is flat & hence the more pressure is acting

on the flat surface which help to increase the lift over

the airfoil. So below fig 7 will give you clear view of





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pressure distribution on the surface of an airfoil at 0°

angle of attack.

Fig 7: Pressure co-efficient on surface of airfoil

Fig 8: Pressure coefficient vs chord length graph

For making it clearer, the effect of fluid over the

airfoil is explained by velocity vector. Fluid is always

having some impact on the airfoil, from below fig 9 it

can be seen that near the leading edges of an airfoil is

experiencing the high impact of fluid.

Fig 9: Velocity vector on airfoil

X. CONCLUSIONS

This paper has presented a viscous analysis method

suitable for incompressible and low Reynolds numberairfoils. The Realizable k- ε solution is been used for

analysis helps to converge our results. The Boussinesqrelationship has helped us to predict the lift & drag

coefficient of an airfoil. The results show that the

present analysis method can accurately predict airfoil

performance at low Reynolds numbers.

XI. REFERENCES

[1] Introduction to flight by John D Anderson

[2] BY VANCE A. TUCKER AND G. CHRISTIAN

PARROTT, AERODYNAMICS OF GLIDING

FLIGHT IN A FALCON AND OTHER BIRDS,

Duke University, Durham, North Carolina 26

September 1969[3] Steven D. Miller, Lift, Drag & Moment for a

NACA-0015 airfoil, The OHIO state university,

28 MAY 2008

[4] 10. Torres, G. E., Aerodynamics of Low Aspect

Ratio Wings at Low Reynolds Numbers withApplications to Micro Air Vehicle Design,

University of Notre Dame, Notre Dame, Indiana,

2002.

[5] Model aircraft Aerodynamics by Martin Simsons

(page -64-65)

airfoil selection of mav (miniature air vehicle)

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