# Lab Chordwise Pressure Distribution Engineering Essay

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This experiment investigates the pressure distribution around a 2D NACA 23015 aerofoil with a flap having a 30-percent chord angle of attack up to about 24o at various of incidences and flap deflections. This experiment was conducted in the 0.6 m x 0.6 m Open Return Low Speed Tunnels at Reynolds Number of 400,000, based on the air speed of 25 m/s and the 0.25m aerofoil chord. The data shows that the flapped aerofoil reduces the adverse pressure gradients and the tendency of the main aerofoil to stall. The flap also influences the air flow around the main aerofoil so that the aerofoil carries a much greater load without stalling. Error between the experimental and theoretical data will give an insight to the limitations of various assumptions such as boundary conditions, wind tunnels, infinite wing.

## Introduction

The flow velocity is modified when it flows over an aerofoil. Using an inviscid fluid model, it is possible to calculate the chordwise pressure distribution on the surface of the aerofoil. The pressure coefficient (defined as where q = dynamic pressure =) has mostly negative values, this indicates the flow accelerates over the upper surface of the aerofoil and the surface static pressure is less than freestream. Near the leading edge on the upper surface, normally there is a large suction peak followed by a region of increasing static pressure (adverse pressure gradient) to the trailing edge. In addition, at the stagnation point near the leading edge Cp has a value of 1.0. The area between the pressure coefficient distributions against x/c gives a good indication of the total lift coefficient or normal force coefficient on the aerofoil.

The suction peak on the upper surface develops and the adverse pressure gradient turns out to be larger when the angle of attack is increased. At the critical angle of attack, the adverse pressure gradient becomes larger and larger and this leads to the boundary layer to separate from the upper surface of the aerofoil. The subsequent result is stall, and the pressure distribution curves will fall on both surfaces (i.e. upper and lower).

In general, a flap at the trailing edge of an aerofoil appears to be one of the most satisfactory high-lift devices to increase aircraft performance. It is capable of developing high lift coefficients and that it gives lower drag at these high lift coefficients. The flap extends past the trailing edge of the wing, thereby resulting in an increased over-all chord and wing area. The principal purpose of this investigation is to determine the effects of wing incidence, the effects of flap deployment, and pressure distributions around the aerofoil at several of angle of attack and at different flap angles.

The coefficient of lift, drag and moment are a function of Reynolds number, and it is the ratio of inertial forces to viscous forces, calculate using the equation (1.1), below.

\mathrm{Re} = {{\rho {\mathbf v} L} \over {\mu}} = {{{\mathbf v} L} \over {\nu}} (Equation 1.1)

where:

{\mathbf v} is the mean velocity of the object relative to the fluid (SI units: m/s)

{L} is a characteristic linear dimension, (travelled length of the fluid; hydraulic diameter when dealing with river systems) (m)

{\mu} is the dynamic viscosity of the fluid (PaÂ·s or NÂ·s/mÂ² or kg/(mÂ·s))

{\bold \nu} is the kinematic viscosity ( {\bold \nu} = \mu /{\rho} ) (mÂ²/s)

{\rho}\, is the density of the fluid (kg/mÂ³).

When the velocity approaches that of sound in the fluid, the lift and drag coefficient becomes a function of Mach number, that is, the ratio of the fluid velocity (relative to the body) to the velocity of sound in the fluid. Calculate using the equation (1.2), below.

\ M = \frac {{v}}{{a}} (Equation 1.2)

where:

\ M is the Mach number,

\ v is the velocity of the source relative to the medium and

\ a is the speed of sound in the medium.

## 2 Method

The current experiment was conducted in the 0.6m by 0.6m Open Return Low Speed Tunnels. There are honeycombs to ensure the flow in the test section stays as uniform as possible.

The model consisted of an NACA 23015 wing with a chord of 0.25m with 30% hinged flap. A row of static pressure tappings on the upper and lower surfaces of both the main aerofoil and flap are placed at mid-span section. The position of the tappings is shown in Appendix A. Pressures are measured by an electric pressure transducer, which is connected to each pressure tapping connection in turn by a scanivalve. This is controlled by a computer, which also logs the data, and presents the pressure distribution directly as a plot of Cp against x/c. For each pressure distribution, the coefficients of lift, pressure drag, pitching moment and flap hinge moment are calculated, and the results are presented in tabular and graphical form.

The average test Reynolds number, based on the flow velocity of 25m/s and 0.25m chord, is approximately 400,000. The test was conducted under standard sea level condition. The Mach number (using Equation 1.2) is M = v/a = 25/340.3 = 0.07 << 0.3, therefore compressible effects can be neglected.

The main portion of the investigation consisted in determinations of lift, drag, and pitching moment for flap deflections of -10o, 0o, 10o, 20o, 40o and 55o throughout an angle-of-attack range from -5o to beyond the stall for the aerofoil.

## 3 Results

## Figure 3.1. - Coefficient of Pressure against x axis at 5o incidence

## Figure 3.2. - Coefficient of Pressure against x axis at 10o incidence

## Figure 3.3. - Coefficient of Pressure against x axis at 20o incidence

## Figure 3.4. - Coefficient of Lift against Angle of Attack

## Figure 3.5. - Coefficient of Drag against Angle of Attack

## Figure 3.6. - Coefficient of Moment against Angle of Attack

## Figure 3.7. - Coefficient of Hinge Moment against Angle of Attack

## Figure 3.8. - Coefficient of Lift against Angle of Attack (Experimental vs Theoretical)

## 4 Discussion

Comparison of pressure diagrams for the aerofoil with flap at the same angle of attack (fig. 3.1. and fig. 3.2.) shows that the flap increases the negative pressure over the entire upper surface of the main aerofoil and increases the positive pressure on the lower surface near the trailing edge. The pressure gradients remain about the same except at the trailing edge of the main aerofoil, where they are reduced. The pressures on the upper and the lower surfaces of the flap both increase with flap deflection. The important effect of the flap in this case is its ability to influence the air flow around the main aerofoil so that the aerofoil carries a much greater load without stalling than is possible without the flap.

For the angle of attack at 20o with flap deflection of 40o (fig. 3.3.) demonstrates that the magnitudes of the peak pressures at the leading edge of the main aerofoil are reduced and the magnitudes of both positive and negative pressures at the trailing edge of the main aerofoil and at the leading edge of the flap are increased. One of the reason which may cause this stall is the air speed is too low. In addition, to obstructing the flow of air below the aerofoil and causes the pressures to build up on the lower surface, the flap influences the air flowing through the slot over the upper surface of the flap to produce a higher average velocity and increases the negative pressure on the flap upper surface. Thus, the influence of the flap is to reduce the adverse pressure gradients and the tendency of the main aerofoil to stall.

By increasing the flap deflection in the camber line effectively alters the camber so that the contribution due to flap deflection is the effect of an additional camber-line shape. The resulting effect is increased effective camber of the wing section and increased in wing area therefore the flap is able to increase the maximum lift of a wing (fig. 3.4.), at the same time reducing the stall angle and increasing drag at a given angle of attack (AoA)(fig. 3.5.). The stall AoA usually will be lower for a wing with deflected flaps than a wing without flaps. This is due to the fact that the pressure gradients at the CL max for both cases are roughly equal. As the wing shape changes as a result of flap deflection, all or some of the components of skin friction drag D0 will increase, and the induced drag Di will also increase due to the change in spanwise lift distribution. The deployment of flap is very useful during final approach down to landing as this would allow a low speed steep descent.

The flow over an aerofoil is two-dimensional. A finite wing is a three-dimensional. On the wing there is a high pressure on the bottom surface and a low pressure on the top surface. The wing on an airplane experiences a much higher pressure drag than an airfoil due to the adverse aerodynamic effects of the wing tips. The difference in pressure at the wing top creates some vortices downstream of the wing which induce a small downward component of air velocity in the neighborhood of the wing itself. This is known as downwash. There is a drag created by the presence of downwash. This additional pressure drag is called induced drag. This component of drag cannot be clearly identify as it is part of the pressure drag. Resolution of normal and chordwise pressure forces in directions normal to and along the relative wind direction give lift and form drag forces. In this case where only pressure forces are considered, the drag coefficient will be the form drag coefficient, which does not include the contribution from skin friction.

## Figure 4.1. - Regions of adverse and favourable pressure gradient

From A to C, when the pressure decreases (and, correspondingly, the velocity along the edge of the boundary layer increases) with passage along the surface the external pressure gradient âˆ‚p/âˆ‚x is negative. Such a pressure gradient is said to be favourable.

Beyond C, the pressure increases and mainstream velocity decreases along the surface. The external pressure gradient is now said to be unfavourable or adverse. The fluid, beyond C, has less momentum than fluid further out, and so when its momentum is reduced still more by the net pressure force the fluid near the surface is soon brought to a standstill. The value of âˆ‚u/âˆ‚y at the surface is then zero as at D. If the adverse pressure gradient is sufficiently strong or prolonged, the flow near the wall is so greatly decelerated that it begins to reverse direction. At E, flow reversal indicates that the boundary layer has separated from the surface. Separation is caused by the reduction of velocity in the boundary layer, adverse pressure combined with a positive pressure gradient (known as an adverse pressure gradient since it opposes the flow). Separation can therefore occur only when an adverse pressure gradient exists.

At low values of Re this may permit a laminar boundary layer to extend into the adverse pressure gradient region of the aerofoil. As a laminar boundary layer is much less able than a turbulent boundary layer to overcome an adverse pressure gradient, the flow will separate from the surface at a lower angle of incidence. This causes a reduction of CLmax. This is a problem that exists in model testing when it is always difficult to match full-scale and model Reynolds numbers. As seen on figure 3.1., the boundary layer transition takes place at about x/c = 0.6. This is caused by the separation bubble, where the pressure gradient approaches zero. Several factors could influence transition from laminar to turbulent flow: Increase surface roughness, increase turbulence in the freestream, adverse pressure gradients, Mach number and heating of the fluid by the surface.[1]

As seen on figure 3.8., the flow first separate from the aerofoil at the angle of attack of 20o. Due to its gradual or docile stall quality this would yield a safest flight. This would be suitable for aircrafts which are non-manoeuvrable.

The lift curve slope for zero flap deflection could be found using

As the experimental wing is not perfectly infinite, this means the aspect ratio is small whereas a 2D aerofoil assumes infinite aspect ratio. As a result, the lift curve slope reduces as aspect ratio reduces. This is due to the downwash angle produced, which reduces the effective angle of attack of the wing.

The aerodynamic center is the reference point about which the aerodynamic moment does not change with changes in angle-of-attack:

(Equation 4.1)

The location of the aerodynamic center can be determined from experimental data from its definition:

(Equation 4.2)

(Equation 4.3)

Using Equation 4.3, for zero flap deflection. The quarter-chord point is the theoretical location of the aerodynamic centre for a camber aerofoil.

Flap Setting

hac

0

0.259

10

0.253

20

0.246311

40

0.253882

55

0.253891

## Table 4.1. - Flap setting and position of aerodynamic centre

Projected frontal area of the aerofoil at zero incidence = 0.25m*15%*0.5m = 0.01875m2. The cross sectional area of the tunnel = 0.6m*0.6m = 0.36m2. Thus, the ratio = 5.21% for zero incidence. It has long been a standard for low-speed wind tunnel testing to operate within an area-ratio of (tunnel cross-section to swept area of a model) 1-10%, proposed by Pope and Harper, (1966) in their text "Low-Speed Wind Tunnel Testing" and earlier by Pankhurst and Holder (1952) in their text "Wind-Tunnel Technique: An Account of Experimental Methods in Low- and High-Speed Wind Tunnels". The open test-section or open jet type of wind tunnel has the capability to allow the conditions inside the test section to be largely unaffected by larger blockage percentage static models because of the ability to leak ï¬‚ow and expand the ï¬‚ow around objects within the test-section. Because of the ability to allow the ï¬‚ow to expand, models can generally be allowed to exhibit higher blockage percentage in open type testing.

15) How will flap deployment change the handling of an aircraft?

A flapped-aerofoil characteristic that is of great importance in stability and control

calculations, is the aerodynamic moment about the hinge line.

The deflection of the flap about a hinge in the camber line effectively alters the camber so that the contribution due to flap deflection is the effect of an additional camber-line shape.

The pressure tabs are placed in the mid span of the aerofoil to ensure only two dimensional flow is considered and any induced drag and vortex drag and sidewall boundary layer could be ignored.

The effective Reynolds Number takes account of the turbulence in the air stream (i.e. Effective Reynolds Number = test Reynolds Number x turbulence factor)

14) How do the wind tunnel walls influence the results?

## Advantages of the Open Return Tunnel

Low construction cost.

Superior design for propulsion and smoke visualization. There is no accumulation of exhaust products in an open tunnel.

## Disadvantages of the Open Return Tunnel

Poor flow quality possible in the test section. Flow turning the corner into the bellmouth may require extensive screens or flow straighteners. The tunnel should also be kept away from objects in the room (walls, desks, people ...)that produce asymmetries to the bellmouth. Tunnels open to the atmosphere are also affected by winds and weather.

High operating costs. The fan must continually accelerate flow through the tunnel.

Noisy operation. Loud noise from the fan may limit times of operation.