aerodynamic analysis on naca 0018 aerofoil
TRANSCRIPT
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]);