using fiosim3 to simulate airfoil - student project

SIM UNIVERSITYSCHOOL OF SCIENCE AND TECHNOLOGY

EAS301e – Aerospace Dynamics

Lab Assignments – 1A

Name Student IDYong Heng Meng H1270977

EAS 301e – AEROSPACE DYNAMICS

1.0 Aim of Report

The aim of this report is to use the undergraduate version of FoilSim III program to:

i. Obtain data that show how lift and drag change with speed, density and wing area respectively. Thereafter, investigate the relationship between lift and speed, with different density and wing area using a scatter plot provided by MS Excel as an aid.

ii. Show that the results are consistent with the definition of lift and drag coefficient from the data obtained in part i stated above.

iii. Obtain data to show how the lift coefficient changes with factors affecting the lift coefficient of an airfoil.

2.0 Scope of Setup

The setup for this laboratory experimental uses the virtual wind tunnel simulated by the undergraduate version of FoilSim III program which is accessible at:

http://www.grc.nasa.gov/www/k12/airplane/foil3u.html

The FoilSIM III analysis will be based on the default lift calculation includes a Stall Model that decreases the lift for angles of attack greater than 10 degrees. The option is to use the Ideal Flow analysis that neglects viscous and compressibility effects. The actual calculations are done with ideal flow and the stall model corrects for flow separation at high angles of attack. The actual calculation is done for a two-dimensional foil.

However, for the analysis of a 2D airfoil for this lab experiment, the following Aspect Ratio Correction and Induced Drag will be set to Off position which will exclude the 3D wing tip effects on lift and the 3D wing tips effects on drag respectively.

3.0 Setup of FoilSim III

FoilSim III can be regarded as a virtual wind tunnel where the investigator can control flight parameters, shape of the airfoil being tested, and the size of the wing being tested – using the flight, shape and size buttons in the control window.

i. Starting with selecting Metric from the Units dropdown list in the control window.

Figure 1 - Metric from the Units dropdown list in the control window


ii. Click on Analysis in the control window, and select AR Off and ID Off.

Figure 2 - AR Off and ID Off from the Analysis in the control window

iii. Starting with the default settings for Shape as shown in Figuree below, and using “Earth – Average Day” conditions for Flight in the control window.

Figure 3 - Default Shape and “Earth – Average Day” conditions for Flight in the control window

iv. Click on Flight in the control window, and enter 100 in the Speed - km/h field follows by an increment of 20 km/h steps till 400 km/h and obtain the Lift and Drag results in the control window for each increment speed entered with density fixed at 1.224 kg/m3 and wing area fixed at 9.290304 m2. The obtained data is presented in later section - Presentation of Findings.

Figure 4 - Enter value for Speed - km/h field and obtain the Lift and Drag results

v. Reset FoilSim and set the settings stated in steps i to iii. Click on Flight in the control window, and fixed the Speed - km/h field as 200. Enter the Altitude-m field with the value of 0 follows by an increment of 1000m steps till 15000m and obtain the Lift and Drag results in the control window for each increment speed entered with speed fixed


at 200 km/h and wing size area fixed at 9.290304 m2. The obtained data is presented in later section - Presentation of Findings.

Figure 5 - Enter value for Altitude-m field with fixed speed, wing size area and obtain the Lift and Drag results

vi. Reset FoilSim and set the settings stated in steps i to iii. Click on Flight in the control window, and fixed the Speed - km/h field as 200 and the Altitude-m as 0 . Using the default value of Wing Size Area - m2 field as 9.290304 follows by an increment of 10 m2 steps till 160 m2 and obtain the Lift and Drag results in the control window for each increment wing size area entered with speed fixed at 200 km/h and altitude fixed at 0 m (density fixed at 1.224 kg/m3). The obtained data is presented in later section - Presentation of Findings.

Figure 6 - Enter value for Wing Size Area - m 2 field with fixed speed, altitude to obtain the Lift and Drag

vii. Reset FoilSim and set the settings stated in steps i to iii. Click on Shape in the control window, and fixed the Speed - km/h field as 200 and the Altitude-m as 0 . Set value Angle-deg field as 0 degree follows by an increment of 1 degree step till 20 degree and obtain the Lift Coefficient result in the control window for each increment wing size area entered with speed fixed at 200 km/h and altitude fixed at 0 m (density fixed at 1.224 kg/m3). The obtained data is presented in later section - Presentation of Findings.


Figure 7 - Enter value for Angle-deg field with fixed speed, altitude to obtain the Lift Coefficient

viii. Reset FoilSim and set the settings stated in steps i to iii. Click on Shape in the control window, and fixed the Speed - km/h field as 200 and the Altitude-m as 0 . Set value Thick-%crd field as 1 % follows by an increment of 1 % step till 20 % and obtain the Lift Coefficient result in the control window for each increment wing size area entered with speed fixed at 200 km/h and altitude fixed at 0 m (density fixed at 1.224 kg/m3). The obtained data is presented in later section - Presentation of Findings.

Figure 8 - Enter value for Thick-%crd field with fixed speed, altitude to obtain the Lift Coefficient

ix. Reset FoilSim and set the settings stated in steps i to iii. Click on Shape in the control window, and fixed the Speed - km/h field as 200 and the Altitude-m as 0 . Set value Camber-%crd field as 0 % follows by an increment of 1 % step till 19 % and obtain the Lift Coefficient result in the control window for each increment wing size area entered with speed fixed at 200 km/h and altitude fixed at 0 m (density fixed at 1.224 kg/m3). The obtained data is presented in later section - Presentation of Findings.

Figure 9 - Enter value for Camber-%crd field with fixed speed, altitude to obtain the Lift Coefficient

x. Reset FoilSim and set the settings stated in steps i to iii. Click on Shape in the control window, and varies the Angle-degree field from 0 degree to 20 degree with 1 degree step each and obtain the Lift Coefficient result in the control window for each basic shapes: Symmetric, High Camber and Flat Plate. The obtained data is presented in later section - Presentation of Findings.


Figure 10 - Enter value for Angle-deg field for Symmetric/High Camber/Flat Plate to obtain the Lift Coefficient

4.0 Presentation of findings

i. Verify the theoretical basis on the lift and drag change with speed used by FoilSim

The first exercise is to examine the simulated Lift and Drag relationship with Speed in FoilSim III. Below assumptions are made for computation purposes:


Chord:

1.524 m Wing Size Area: 9.290304 m2 Density: 1.224 kg/m3

Span: 6.096 m Angle of Attack: 5.0 Airfoil Shape: Symmetric

A value of 100 km/h was entered in the Speed - km/h field follows by an increment of 20 km/h steps till 400 km/h to obtain the Lift and Drag results in the control window of the FioSim III program. A default symmetrical airfoil is used to simulate across speed ranging from 100 to 400 km/h. The simulated results for lift and drag were tabulated in the following table and plot unto the scatter graph:

Table 1 - Lift and drag vs. speed


Altitude: 0m (Earth - Average Day)FoilSim Results Calculated Results

Speed (km/h)

V (m/s)

Lift (N)

Drag (N)

Lift Coeff.

Cl

Drag Coeff.

Cd

100 27.778 2685 91 0.611 0.020 2680.51 87.74120 33.333 3866 128 0.611 0.020 3859.94 126.35140 38.889 5262 172 0.611 0.020 5253.80 171.97160 44.444 6874 221 0.611 0.020 6862.11 224.62180 50.000 8700 277 0.611 0.019 8684.85 270.07200 55.556 10741 338 0.611 0.019 10722.04 333.42220 61.111 12996 404 0.611 0.019 12973.67 403.44240 66.667 15467 477 0.611 0.018 15439.74 454.85260 72.222 18152 555 0.611 0.018 18120.25 533.82280 77.778 21052 638 0.611 0.018 21015.20 619.11300 83.333 24167 727 0.611 0.018 24124.60 710.71320 88.889 27497 821 0.611 0.018 27448.43 808.63340 94.444 31042 921 0.611 0.018 30986.70 912.87360 100.000 34801 1026 0.611 0.018 34739.42 1023.42380 105.556 38775 1137 0.611 0.017 38706.58 1076.94400 111.111 42965 1253 0.611 0.017 42888.17 1193.29

SCVL L2

21 SCVD D

2

21

50 100 150 200 250 300 350 400 4500

5000

1000015000

2000025000

30000

3500040000

45000

0

500

10001500

20002500

3000

3500

40004500

Lift and Drag vs Speed Lift (N) Drag (N)

Speed (km/h)

Lift (N

) Drag (N)

Figure 11 - Scatter Plot of lift and drag vs. speed

Analysis:

From the data obtained from FoilSim III, the curves for Lift and Drag are of parabolic type (quadratic function). This is because the factor of Lift and Drag are the square of speed.

The relationship between the two can be summarized as below:

L=12

ρV 2CL S

D=12

ρV 2 CD S

ii. Verify the theoretical basis on the lift and drag change with density used by FoilSim


Where S is the Area of the Airfoil,

V is the Altitude and

ρ is the density of air.

The first exercise is to examine the simulated Lift and Drag relationship with Density in FoilSim III. Below assumptions are made for computation purposes:

Chord:

1.524 m Wing Size Area: 9.290304 m2 Speed: 200 km/h

Span: 6.096 m Angle of Attack: 5.0 Airfoil Shape: Symmetric

Fixed the Speed - km/h field as 200 km/h. With the Altitude-m set with the value of 0m follows by an increment of 1000m steps till 15000m to obtain the Lift and Drag results in the control window of the FioSim III program. A default symmetrical airfoil is used to simulate across altitude ranging from 0 to 15000 m. The simulated results for lift and drag were tabulated in the following table and plot unto the scatter graph:

Altitude: 0m, Speed: 200 km/h, Varies density FoilSim Results Calculated Results

Speed (km/h)

Lift(N)

Drag(N)

Altitude(m)

density, ρ

(kg/m3)

Lift Coeff.

Cl

Drag Coeff.

Cd

200 10741 338 0 1.224 0.611 0.019 10722.04 333.42200 9748 309 1000 1.110 0.611 0.019 9723.42 302.36200 8828 282 2000 1.006 0.611 0.019 8812.40 274.04200 7975 257 3000 0.908 0.611 0.019 7953.93 247.34200 7187 234 4000 0.819 0.611 0.019 7174.31 223.10200 6460 212 5000 0.736 0.611 0.020 6447.24 211.04200 5790 192 6000 0.659 0.611 0.020 5772.73 188.96200 5175 173 7000 0.589 0.611 0.020 5159.55 168.89200 4611 156 8000 0.525 0.611 0.020 4598.92 150.54200 4096 140 9000 0.466 0.611 0.020 4082.09 133.62200 3625 125 10000 0.413 0.611 0.021 3617.81 124.34200 3198 112 11000 0.364 0.611 0.021 3188.58 109.59200 2716 96 12000 0.309 0.611 0.021 2706.79 93.03200 2319 84 13000 0.264 0.611 0.022 2312.60 83.27200 1981 73 14000 0.225 0.611 0.022 1970.96 70.97200 1692 63 15000 0.192 0.611 0.023 1681.89 63.31

Table 2 - Lift and drag vs. density


SCVL L2

21 SCVD D

2

21

0.000 2.000 4.000 6.000 8.000 10.000 12.000300

2300

4300

6300

8300

10300

50150250350450550650750850950

Lift and Drag vs Density Lift (N) Drag (N)

Density, ρ (kg/m3)

Lift (N

)Drag (N

)

Figure 12 - Scatter Plot of lift and drag vs. density

Analysis:

From the data obtained from FoilSim III, the graphs above shows the lines are relatively straight. This concludes that both lift and drag varies linearly with changes in density. Increase in fluid density will increase the drag force and lift force of an airfoil. This is because the Lift and Drag are directly proportional to the density. The relationship between the two can be summarized as below:

L=12

ρV 2CL S

D=12

ρV 2CD S





iii. Verify the theoretical basis on the lift and drag change with wing area used by FoilSim

The first exercise is to examine the simulated Lift and Drag relationship with Wing Area in FoilSim III. Below assumptions are made for computation purposes:

Chord:

1.524 m Wing Size Area: 9.290304 m2 Speed: 200 km/h

Span: 6.096 m Angle of Attack: 5.0 Altitude: 0m

Fixed the Speed - km/h field as 200km/h and the Altitude-m as 0m . Using the default value of Wing Size Area - m2 field as 9.290304 follows by an increment of 10 m2 steps till 160 m2

and obtain the Lift and Drag results in the control window of the FioSim III program. A default symmetrical airfoil is used to simulate across altitude ranging from 9.290304 to 160 m2. The simulated results for lift and drag were tabulated in the following table and plot unto the scatter graph:

Altitude: 0m, Speed: 200 km/h, Varies Wing Area FoilSim Results Calculated Results

Speed (km/h)

Lift(N)

Drag(N)

Area, S (m2)

Lift Coeff.Cl

Drag Coeff.

Cd

200 10741 338 9.290304 0.611 0.019 10722.04 333.42200 23123 704 20 0.611 0.018 23082.22 680.00200 34685 1033 30 0.611 0.018 34623.33 1020.00200 46247 1355 40 0.611 0.017 46164.44 1284.44200 57808 1674 50 0.611 0.017 57705.56 1605.56200 69370 1989 60 0.611 0.017 69246.67 1926.67200 80932 2300 70 0.611 0.017 80787.78 2247.78200 92494 2610 80 0.611 0.017 92328.89 2568.89200 104056 2917 90 0.611 0.017 103870.00 2890.00200 115617 3223 100 0.611 0.017 115411.11 3211.11200 127179 3527 110 0.611 0.016 126952.22 3324.44200 138741 3829 120 0.611 0.016 138493.33 3626.67200 150303 4130 130 0.611 0.016 150034.44 3928.89200 161865 4429 140 0.611 0.016 161575.56 4231.11200 173426 4728 150 0.611 0.016 173116.67 4533.33200 184988 5025 160 0.611 0.016 184657.78 4835.56

Table 3 - Lift and drag vs. wing area


SCVL L2

21 SCVD D

2

21

0 2 4 6 8 10 12 14 16 180

2

4

6

8

10

12

02000400060008000100001200014000160001800020000

Lift and Drag vs Wing Size AreaLift(N)

Drag(N)

Area, S (m^2)

Lift )N

) Drag (N)

Figure 13 - Scatter Plot of lift and drag vs. wing area

Analysis:

From the data obtained from FoilSim III, the graphs above shows the lines are relatively straight. This concludes that both lift and drag varies linearly with changes in wing area. Increase in wing area will increase the drag force and lift force of an airfoil. This is because the Lift and Drag are directly proportional to the wing area. The relationship between the two can be summarized as below:

L=12

ρV 2CL S

D=12

ρV 2CD S





iv. Verify the results are consistent with the definition of lift and drag coefficient used by FoilSim and theoretical basis

Altitude: 0m (Earth - Average Day) FoilSim Results Calculated Results

Speed (km/h)

Lift (N)

Drag (N)

density, ρ

(kg/m3)Area, S

(m2)

Lift Coeff.

Cl

Drag Coeff.

Cd

100 2685 91 1.2249.29030

4 0.611 0.020 0.612 0.021

120 3866 128 1.2249.29030

4 0.611 0.020 0.612 0.020

140 5262 172 1.2249.29030

4 0.611 0.020 0.612 0.020

160 6874 221 1.2249.29030

4 0.611 0.020 0.612 0.020

180 8700 277 1.2249.29030

4 0.611 0.019 0.612 0.019

200 10741 338 1.2249.29030

4 0.611 0.019 0.612 0.019

220 12996 404 1.2249.29030

4 0.611 0.019 0.612 0.019

240 15467 477 1.2249.29030

4 0.611 0.018 0.612 0.019

260 18152 555 1.2249.29030

4 0.611 0.018 0.612 0.019

280 21052 638 1.2249.29030

4 0.611 0.018 0.612 0.019

300 24167 727 1.2249.29030

4 0.611 0.018 0.612 0.018

320 27497 821 1.2249.29030

4 0.611 0.018 0.612 0.018

340 31042 921 1.2249.29030

4 0.611 0.018 0.612 0.018

360 34801 1026 1.2249.29030

4 0.611 0.018 0.612 0.018

380 38775 1137 1.2249.29030

4 0.611 0.017 0.612 0.018

400 42965 1253 1.2249.29030

4 0.611 0.017 0.612 0.018

Table 4 - FoiSim III's Lift and Drag Coefficients are consistent with theoretical basis

Analysis:

From the data obtained from FoilSim III, the Lift and Drag Coefficients are consistent with the definition of lift and drag coefficient used in theoretical basis. Where the relationship between the two can be summarized as below:


SV

LCL2

21

SV

DCD2

21

SCVL L2

21

SCVD D2

21

CL=L12 ρV 2 S

CD=D12

ρV 2S

v. Obtain data to show how the lift coefficient changes with factors affecting the lift coefficient of an airfoil.

a. Lift Coefficient versus angle of attack (AOA): The following data obtained from FioSim III shows that factor like angle of attack is affecting the lift coefficient of an airfoil.





L is the Lift force.

D is the drag force.

Altitude: 0m, Speed: 200 km/h, Area: 9.290304 m^2, 12.5% crd

FoilSim Results Calculated Results

Speed (km/h)

Lift(N)

Drag(N)

AOA (degree

)

density, ρ

(kg/m3)Area, S (m2)

Lift Coeff.

Cl

Drag Coeff.

Cd

200 0 287 0 1.2249.290

3 0.000 0.016 0.000 0.016

200 2150 303 1 1.2249.290

3 0.122 0.017 0.123 0.017

200 4301 315 2 1.2249.290

3 0.244 0.017 0.245 0.018

200 6449 324 3 1.2249.290

3 0.367 0.018 0.367 0.018

200 8596 331 4 1.2249.290

3 0.489 0.018 0.490 0.019

200 10741 338 5 1.2249.290

3 0.611 0.019 0.612 0.019

200 12882 348 6 1.2249.290

3 0.733 0.019 0.734 0.020200 15019 367 7 1.224 9.290 0.855 0.020 0.856 0.021


SV

LCL2

21

SV

DCD2

21

3

200 17151 398 8 1.2249.290

3 0.976 0.022 0.977 0.023

200 19275 448 9 1.2249.290

3 1.098 0.025 1.098 0.026

200 21400 520 10 1.2249.290

3 1.218 0.029 1.219 0.030

200 23398 620 11 1.2249.290

3 1.332 0.035 1.333 0.035

200 25111 752 12 1.2249.290

3 1.430 0.042 1.431 0.043

200 26475 918 13 1.2249.290

3 1.508 0.052 1.509 0.052

200 27429 1118 14 1.2249.290

3 1.562 0.063 1.563 0.064

200 27910 1350 15 1.2249.290

3 1.589 0.076 1.590 0.077

200 27855 1609 16 1.2249.290

3 1.586 0.091 1.587 0.092

200 27204 1886 17 1.2249.290

3 1.549 0.107 1.550 0.107

200 25897 2166 18 1.2249.290

3 1.475 0.123 1.476 0.123

200 23873 2430 19 1.2249.290

3 1.359 0.138 1.360 0.138

200 21075 2652 20 1.2249.290

3 1.200 0.151 1.201 0.151

Table 5 - Lift Coefficient vs. angle of attack

0 5 10 15 20 250.0000.2000.4000.6000.8001.0001.2001.4001.6001.800

Lift Coeff.CL vs AOA

Lift Coeff.Cl

AOA (degree)

Lift C

oeff.

CL


Figure 14 - Scatter Plot of lift coefficient vs. angle of attack (AOA)

Analysis:

The plot above shows how the lift coefficient varies with angle of attack for a symmetric airfoil. At low angles, the lift coefficient is nearly linear. Notice on this plot that at zero angle of attack, the lift is zero as the airfoil is symmetric in shape used in this analysis. The lift coefficient can be determined mathematically. For example: for thin airfoils at

subsonic speed, and small angle of attack, the lift coefficient CL is given by: CL=2Παwhere pi is 3.1415, and α is the angle of attack expressed in radians: pi radians = 180 degrees.

b. Lift Coefficient versus airfoil thickness: The following data obtained from FioSim III shows that factor like airfoil thickness is affecting the lift coefficient of an airfoil.

Altitude: 0m, Speed: 200 km/hFoilSim Results

Calculated Results


Speed (km/h)

Lift(N)

Drag(N)

Thick

(% crd)

Lift Coeff.

Cl

Drag Coeff.

Cd

200 9709 377 1 0.553 0.021 0.553 0.021200 9804 377 2 0.558 0.021 0.559 0.021200 9898 377 3 0.563 0.021 0.564 0.021200 9991 377 4 0.569 0.021 0.569 0.021200 10082 377 5 0.574 0.021 0.575 0.021200 10173 352 6 0.579 0.020 0.580 0.020200 10263 327 7 0.584 0.018 0.585 0.019200 10352 302 8 0.589 0.017 0.590 0.017200 10440 277 9 0.594 0.015 0.595 0.016200 10527 252 10 0.599 0.014 0.600 0.014200 10613 286 11 0.604 0.016 0.605 0.016200 10699 321 12 0.609 0.018 0.610 0.018200 10783 355 13 0.614 0.020 0.614 0.020200 10866 389 14 0.618 0.022 0.619 0.022200 10948 423 15 0.623 0.024 0.624 0.024200 11030 426 16 0.628 0.024 0.629 0.024200 11110 429 17 0.623 0.024 0.633 0.024200 11190 433 18 0.637 0.024 0.638 0.025200 11269 436 19 0.641 0.024 0.642 0.025200 11346 439 20 0.646 0.025 0.647 0.025


SV

LCL2

21

SV

DCD2

21

Table 6 - Lift Coefficient vs. airfoil thickness

Analysis:

The table above shows how the lift coefficient varies with airfoil thickness for a symmetric airfoil. The table shows a linear relationship between airfoil thickness and lift coefficient.

• Influence of airfoil thickness on lift coefficient– CLmax increases with increasing airfoil thickness, up to airfoil thickness in the

order of approximate 13-16% and at airfoil thickness > 16% begins to CLmax

decrease.• Influence of airfoil thickness on drag coefficient

– CDmin increases with increasing airfoil thickness.

c. Lift Coefficient versus camber: The following data obtained from FioSim III shows that factor like camber is affecting the lift coefficient of an airfoil.



Altitude: 0m, Speed: 200 km/h, Area: 9.290304 m^2, 5 deg AOA

FoilSim Results

Calculated Results

Speed (km/h)

Lift(N)

Drag(N)

Camber

(% c)

Lift Coeff.

Cl

Drag Coeff.

Cd

200 10740 377 0 0.611 0.019 0.612 0.021200 12908 381 1 0.735 0.021 0.736 0.022200 15076 424 2 0.858 0.024 0.859 0.024200 17245 467 3 0.982 0.026 0.983 0.027200 19413 510 4 1.105 0.029 1.106 0.029200 21582 553 5 1.229 0.031 1.230 0.032200 23750 618 6 1.352 0.035 1.353 0.035200 25919 682 7 1.476 0.038 1.477 0.039200 28088 747 8 1.599 0.042 1.601 0.043200 30256 811 9 1.723 0.046 1.724 0.046200 32424 876 10 1.846 0.049 1.848 0.050200 34593 1027 11 1.970 0.058 1.971 0.059200 36760 1178 12 2.093 0.067 2.095 0.067200 38928 1329 13 2.217 0.075 2.218 0.076200 41095 1481 14 2.340 0.084 2.342 0.084200 43262 1632 15 2.464 0.092 2.465 0.093200 45428 1722 16 2.587 0.098 2.589 0.098200 47594 1812 17 2.710 0.103 2.712 0.103200 49759 1902 18 2.834 0.108 2.836 0.108200 51923 1992 19 2.957 0.113 2.959 0.114

SV

LCL2

21

SV

DCD2

21

Table 6 - Lift Coefficient vs. camber

Analysis:

The table above shows how the lift coefficient varies with camber for a symmetric airfoil. The table shows a linear relationship between airfoil thickness and lift coefficient. An increasing camber increases zero lift angle of attack. Camber can also be used to get a low drag coefficient at the designed lift coefficient.

d. In summary: Lift Coefficient is affected by Angle of Attack (AOA), airfoil thickness and camber. This can be shown in data obtained in FioSim III program by comparing the basic airfoil shape of Symmetric, High Camber and Flat Plate.

Using the following parameters from the FioSim III basic shape:

Symmetric Airfoil Camber: 0 %c Thick-%crd: 12.495 High Camber Airfoil Camber: 15 %c Thick-%crd: 12.5Flat Plate Camber: 0 %c Thick-%crd: 1.0

The obtained Lift Coefficient from FioSim III is shown below:


AOA (degree

)

Symmetrical Airfoil Lift

Coeff.Cl

High Camber Airfoil Lift

Coeff.Cl

Flat Plate Airfoil

Lift Coeff.Cl

0 0.000 1.857 0.0001 0.122 1.980 0.1092 0.244 2.102 0.2193 0.367 2.223 0.3284 0.489 2.344 0.4385 0.611 2.464 0.5476 0.733 2.583 0.6567 0.855 2.701 0.7658 0.976 2.819 0.8749 1.098 2.936 0.98210 1.218 3.052 1.09111 1.332 3.151 1.19212 1.430 3.215 1.28013 1.508 3.241 1.34914 1.562 3.225 1.39815 1.589 3.164 1.42216 1.586 3.055 1.42017 1.549 2.895 1.38618 1.475 2.681 1.32019 1.359 2.409 1.21720 1.200 2.077 1.074

Table 7 - Varies airfoil geometry vs. angle of attack (AOA)

0 5 10 15 20 250.000

0.500

1.000

1.500

2.000

2.500

3.000

3.500

Varies Airfoil Geometry affecting Lift Coeff. CL vs AOA

Symmetrical Airfoil Lift Coeff.ClHigh Camber Airfoil Lift Coeff.Cl Flat Plate Airfoil Lift Coeff.Cl

AOA (degree)

Lift

Coeff

. CL

Figure 15 - Scatter Plot of lift coefficient vs. angle of attack (AOA) for varies airfoil geometry

Analysis:

Lift Coefficient increases with an increase in Angle of Attack (AOA), airfoil thickness and camber. The maximum lift coefficient of moderately cambered airfoil sections increase with increasing camber. However, camber can affect the stall behavior, for airfoil sections


that have the maximum camber far forward the stall is very abrupt. The detail investigation of stall will be discussed in the next lab assignment later.


using fiosim3 to simulate airfoil - student project

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