SCHOOL OF MECHANICAL ENGINEERING

Lift and Drag of an AerofoilMaster of Mechanical Engineering

Aerodynamics Lab Part 1

Session date: 22/10/2014

Session Time: 2pm-5pm

Due Date: 05/11/2014

By

Laveti Tejaswi (a1659176)

&

Lovedeep Singh (a1674234)

Fluid Dynamics Part 1

Lift and Drag of an Aerofoil

AIM:1. To calculate the pressure distribution on a NACA 0018 aerofoil2. Find out the lift, drag coefficients and their ratio’s3. Locate the stall angle and observe the flow for a fully stalled situation4. Compare the results with previously published engineering data5. Determine the quality or accuracy of the results

APPARATUS:1. A closed wind tunnel with an open working section2. A pitot tube which measures the total stagnation pressure3. A NACA 0018 cross section aerofoil with pressure tapping on the top and

bottom4. Multi-tube manometer which helps to find out the pressure readings at the

top and bottom of the aerofoil

Pressure tapping locations on the aerofoil

COMPONENTS OF THE WIND TUNNEL 1. Collector: it is used to collect the flow from the wind tunnel and make a

closed system2. Turning vanes: used to decrease the pressure loss in the flow3. Diffuser: used to slow down the flow and increase the pressure4. Axial fan: can increase or decrease the desired speed5. Wire screens: are used to create a uniform flow and reduce any turbulence or

decay them faster6. Manometer: used to measure the pressure on the top and bottom of the

aerofoil7. Convergent-divergent Section: increase the velocity and decrease the

pressure

NACA 0018 AEROFOIL:

http://www.aerospaceweb.org/question/airfoils/airfoil/airfoil-parts.jpg

PERFECT CONDITIONS FOR LIFT:

http://www.skybrary.aero/images/Aerofoil1.jpg

Low angle of attack, low pressure at the top and high pressure at the bottom, as flow will always travel from high to low this will cause lift in the aerofoil. But these

conditions are desirable only at the time of flight. During touchdown, the opposite of the above mentioned diagram is required so as to induce more drag, reduce the speed at landing.

THEORY:The aerodynamic forces acting on the aerofoil is made up of two components

1. Normal components 2. Chord wise components

The shear stress represents the chord wise force component and the pressure for the normal force component

Lift is always perpendicular to fluid flow in the upward direction and drag is the force provided in the direction of the flow.

Lift or drag is a function of the following parameters

L =f (α, U, a, ρ, u, c, s, S)

Where α = angle of attack, U = velocity of air, ρ = density of air, u= viscosity of air, c= chord of the aerofoil, s = span of the aerofoil, S = shape of the aerofoil

The coefficient of lift or Drag is a function of the following parameter

CL =f (α, M, Re, AR)

But we could neglect the Mach number because in the experiment the flow is incompressible

Re is the Reynolds number and AR is the aspect ratio = s/c

The coefficient of pressure (Cp) is just the ratio of the surface static pressure to the dynamic pressure.

The value of Cp cannot be greater than 1, as in incompressible flow the stagnation pressure is equal to the total pressure.

However, in compressible conditions, the value of Cp could be greater than 1.

The normal and chord component of force acting on the aerofoil can be given by the

Following equations

The lift coefficient can be derived using the following equations as follows

The point or angle of attack at which there is a sudden or gradual decrease in the lift of an aerofoil is called the stall angle. It is during landing when we want more angle of attack so as to reduce the speed and increase the drag.

PROCEDURE:

1. Set the angle of attack at 0 degrees and note down manometer readings for both the top and bottom.

2. Turn on the wind tunnel for the desired conditions3. There are a total of 15 readings for both the top and bottom sections4. For different angle of attacks note the readings of the manometer as stated

in 1 and 2

5. The last reading is noted while the wind tunnel is turned off6. Tabulate the data collected7. Plot the relevant graphs

NUMERICAL PROCEDURE1. The manometer readings are taken at the beginning and end of the

experiment and averaged2. The readings are then scaled by subtracting the static values for various

angles3. The pressure coefficients are calculated by dividing the readings in point 2

and the dynamic pressure from the pitot tube4. The chord wise and drag wise components are then calculated by integration5. The coefficient of lift and drag are calculated

RESULTSThe dynamic pressure can be calculated from the following equation

Where, ρ is the density of fluid, g is the acceleration due to gravity, H is averaged dynamic pressure

By substituting the values the dynamic pressure is about 228.7645 pa

The flow speed and Reynold’s number can be calculated from the following formulas

Where C is the chord length of the aerofoil and u is the dynamic viscosity

By substitution V =18.8984 m/s and Re = 2.2618 e+ 005

X/C vs Cp

Y/C vs Cp

For calculating the chord wise and normal component , the area integral of the function Cp with respect to y/c and x/c used by the trapezoidal rule function of mat lab and the following equations

Normal and chord wise components

Angle of attack Normal Component Chord wise Component2 ° 0.1798 -0.02186 ° 0.3276 0.018610 ° 0.5134 0.055414 ° 1.8976 0.0231

Lift and Drag Coefficient

Angle of attack Lift coefficient Drag Coefficient2 ° 0.1807 -0.01566 ° 0.3236 0.050110 ° 0.4996 0.137914 ° 1.8246 0.7383

DISCUSSIONThe results are compared from the NACA 0018 literature document

It is obvious that the NACA literature document has conducted the experiment for a wide range of angle of attacks compared to only 4 different angles of attacks done in our experiment.

Comparison of Lift Coefficient

Angle of attack Experimental Data NACA DATA TR-647

%error

2 ° 0.1807 0.1143 -58.096 ° 0.3236 0.4182 22.6210 ° 0.4996 0.7170 30.3214 ° 1.8246 1.005 -81.55

Comparison of Drag Coefficient

Angle of attack Experimental Data NACA DATA TR-647

%error

2 ° -0.0156 0.0363 142.96 ° 0.0501 0.0890 43.710 ° 0.1379 0.1799 23.3414 ° 0.7383 0.3611 -104.4

The closest coefficient of lift we got is for an angle of attack of 6 ° and for the Drag Coefficient is at an angle of attack for 10 °. From the graphs the stall angle was occurred at around 17°

SOURCES FOR ERRORS1. Different experimental set up2. Positioning of the aerofoil3. Shape of the aerofoil4. Equipment used5. Human errors6. Surface roughness7. Material of the aerofoil8. Skin friction Coefficient9. Reynold’s number10.Temperature at the time of experiment11.Calibration errors in the instruments

Lift Coefficient Vs angle of attack (NACA DATA (TR-647) vs Experimental DATA)

Drag Coefficient vs angle of attack (NACA DATA (TR-647) vs Experimental DATA)

CONCLUSIONStall occurs at a high angle of attack. Whenever the airplane is about to land, more drag is to be induced to slow down the craft and reduce the lift. The flow separation occurs at the end of the aerofoil. More is the wake; more would be the drag and less lift.

However, errors might have arisen due to several factors but the main idea of the lab was to understand the flow of an aerofoil at various angles of attack and its relation to the drag and lift. It was very interesting when at an angle of attack of 30°; the flow had a different pattern at different locations of the aerofoil. This was clearly understood by tying a thread to a small stick and presented in the flow. The shear stress was not considered due to the complexity.

REFERENCES1. http://classicairshows.com/Education/Aerodynamics/AeroDynamicsImages/

Airfoil2CL.gif 2. Applied aerodynamics – Compressible & incompressible flow, Professor

Kelso,R (MECH ENG School of Mechanical Engineering, The University of Adelaide

3. Munson, BR Young, DF Okiishi, TH (1998), ‘Fundamental of fluid mechanics’, 3rd Edition, Wiley, New York

4. NACA 0018 aerofoil characteristics (Source: NACA TR-647 report)(Goett, HJ et.al 1939)

MATLAB SOURCE CODEA=xlsread(aerolab.xls');P_atm = 1.019*10^5; %Pag = -9.81; %m/s^2u = 1.894 * 10^-5; %kg/msR = 287; %J/kgKT = 299; %Kelvin alpha = [2,6,10,14]; %Degreesa_mono = 78; %Monometer angle in degreesc = 181*10^-3; %Chord Length in m %Extract static condition data points and averageh_1 = A(:,1);h_2 = A(:,6);h_avg =0.5*(h_1+h_2);%Scale the Pressuresh_uncal= A(:,2:5);h_calib= h_uncal - repmat(h_avg,1,4);%Calculating static and dynamic pressuresdhs = h_calib(1:32,:);dhq = h_calib(33,:);%Correcting the Dynamic Pressurerho = P_atm/(R*T);H = mean(dhq);disp('the dynamic pressure is :')q = rho*g*H*sind(90-a_mono)%Calculating the Flow Speeddisp('the velocity of flow is :')v = sqrt(2*q/rho)%Calculating the Reynolds Numberdisp('the reynolds number is :')Re = rho*v*c/u%Calculating the pressure coefficientc_p = dhs./repmat(dhq,32,1);%Load in Pressure Tapping LocationsXonC=xlsread('XonC.xls');YonC=xlsread('YonC.xls'); %Plot the Pressure Coefficents Distributions along the Aerofoilfigureplot(XonC, c_p,'.-')title('The Pressure Coefficent Distribution along the Aerofoil (X on C vs c_p)');xlabel('X on C Value');ylabel('Pressure Coefficent');legend ('2 degree', '6 degrees','10 degrees','14 degrees')figureplot(YonC, c_p,'.-')title('The Pressure Coefficient Distribution along the Aerofoil (Y on C vs

c_p)');xlabel('Y on C Value');ylabel('Pressure Coefficent');legend ('2 degree', '6 degrees','10 degrees','14 degrees')%Calculate the Normal coefficient and the chordwise components of form dragc_n = trapz(XonC,c_p)c_cf = trapz(YonC,c_p)%Calculate the lift and drag coefficientsc_L = c_n.*cosd(alpha)-sind(alpha).*c_cfc_DF = c_n.*sind(alpha)+cosd(alpha).*c_cf%Plot the Lift and Drag coefficients vs angle of Attackfigureplot(alpha, c_L,'.-')title('The Lift Coefficient vs the angle of attack');xlabel('Angle of Attack (Degree)');ylabel('Lift Coefficient');figureplot(alpha, c_DF,'.-')title('The Drag Coefficient vs the angle of attack');xlabel('Angle of Attack (Degree)');ylabel('Drag Coefficient');% extracting the NACA 167 resultsNACA=xlsread('NACA.xls');NACAD=xlsread('NACA DRAG.xls');%compairing Lift dataNACAalp=NACA(:,1);NACACl=NACA(:,2);figure;plot(NACAalp,NACACl,'.-');title('The lift coefficient vs the angle of attack (Comparision with NACA TR-647)');xlabel('angle of attack in degrees');ylabel('Lift Coefficient');hold onexperimental=plot(alpha,Cl,'.--');legend('NACA TR-647','Experimental data');set(experimental,'Color',[1 0 0]);%comparison of DRAGNACADalp=NACAD(:,1);NACADCl=NACAD(:,2);figure;plot(NACADalp,NACADCl,'.-');title('The Drag coefficient vs the angle of attack (Comparision with NACA TR-647)');xlabel('angle of attack in degrees');ylabel('Drag Coefficient');hold onexperimental=plot(alpha,Cd,'.--');legend('NACA TR-647','Experimental data');set(experimental,'Color',[1 0 0]);