Airfoil Selection of MAV (Miniature Air Vehicle) for Low Reynolds Number

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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, Lift 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 will glide 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. Lift to 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 drag coefficient (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 number Re 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 will operate 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

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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 Lift force.

Fig 2 : Level flight condition T=D

L=W

Above 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 the drag 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 a steady (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 the airfoil at different angle of attack [4]. The symmetric airfoil stall at greater angle of attack as compare to cambered 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 is used.

As the airfoil is selected, the gliding of the flight should be good. So airfoil has to be selected in such a way that has less cambered as compared to other airfoils, i.e.: the lower surface of an airfoil has to be nearly 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 any other CAD software. These CAD model is import in

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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 of mesh 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°. Here the in-viscid flow is considered, as air is for MAV incompressible 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 & mass across 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. To understand 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 they will 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 a solid 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 used only 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 of attack for FLUENT. The angle of attack is varied from - 5° to 5° with 1° of interval.

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41 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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42 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 number airfoils. The Realizable k- ε solution is been used for analysis helps to converge our results. The Boussinesq relationship 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 with Applications to Micro Air Vehicle Design, University of Notre Dame, Notre Dame, Indiana, 2002.

[5] Model aircraft Aerodynamics by Martin Simsons (page -64-65)

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